avionic emb 145-1

Business and Commuter Aviation SystemsHoneywell Inc.Box 29000Phoenix, Arizona 85038

PRIMUS® 1000 IntegratedAvionics System

Embraer EMB-145

SystemMaintenance Manual

Volume I ── System Overview and SystemDescription

22-05-14TITLE PAGE T-1

PRINTED IN U.S.A. PUB. NO. A15-1146-065 1 NOVEMBER 1996

SYSTEMMAINTENANCEMANUALEMBRAER 145

PROPRIETARY NOTICE

This document and the information disclosed herein are proprietary data of Honeywell Inc.Neither this document nor the information contained herein shall be used, reproduced, ordisclosed to others without the written authorization of Honeywell Inc., except to the extentrequired for installation or maintenance of recipient’s equipment.

NOTICE - FREEDOM OF INFORMATION ACT (5 USC 552) ANDDISCLOSURE OF CONFIDENTIAL INFORMATION GENERALLY (18 USC 1905)

This document is being furnished in confidence by Honeywell Inc. The information disclosedherein falls within exemption (b) (4) of 5 USC 552 and the prohibitions of 18 USC 1905.

S96

22-05-14TITLE PAGE T-2

Copyright 1996 Honeywell Inc.All Rights Reserved 1 NOVEMBER 1996


RECORD OF REVISIONS - VOLUME I

Upon receipt of a revision, insert the latest revised pages and dispose of superseded pages. Enterrevision number and date, insertion date, and the incorporator’s initials on the Record of Revisions. Thetyped initials HI are used when Honeywell Inc. is the incorporator.

RevisionNumber

RevisionDate

InsertionDate By

RevisionNumber

RevisionDate

InsertionDate By

Page RR-122-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.


(Blank Page)

Page RR-222-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.


LIST OF EFFECTIVE PAGES - VOLUME I

Original . . 0 . . Nov 1/1996

SUBHEADING AND PAGE REVISION SUBHEADING AND PAGE REVISION

TitleT-1 0T-2 0

Record of RevisionsRR-1 0RR-2 0

List of Effective PagesLEP-1 0LEP-2 0LEP-3 0LEP-4 0LEP-5 0LEP-6 0LEP-7 0LEP-8 0

Table of ContentsTC-1 0TC-2 0TC-3 0TC-4 0TC-5 0TC-6 0TC-7 0TC-8 0TC-9 0TC-10 0TC-11 0TC-12 0TC-13 0TC-14 0TC-15 0TC-16 0TC-17 0TC-18 0TC-19 0TC-20 0TC-21 0TC-22 0

TC-23 0TC-24 0TC-25 0TC-26 0TC-27 0TC-28 0TC-29 0TC-30 0

IntroductionINTRO-1 0INTRO-2 0INTRO-3 0INTRO-4 0INTRO-5 0INTRO-6 0INTRO-7 0INTRO-8 0

Table of Contents - Section 1TC1-1 0TC1-2 0

Section 1System Overview

1-1 01-2 01-3 01-4 0

F 1-5/6 0F 1-7/8 0F 1-9/10 0F 1-11/12 0

1-13 01-14 01-15 01-16 01-17 01-18 01-19 01-20 0

F indicates right foldout page with blank back.

Page LEP-122-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.



Section 1 (cont)System Overview

1-21 01-22 01-23 01-24 0

F 1-25/26 0F 1-27/28 0

1-29 01-30 01-31 01-32 01-33 01-34 01-35 01-36 01-37 01-38 0

Section 2System Description

2-1 02-2 0

Table of Contents - Section 2.1TC2-1-1 0TC2-1-2 0TC2-1-3 0TC2-1-4 0TC2-1-5 0TC2-1-6 0TC2-1-7 0TC2-1-8 0

Section 2.1Electronic Display System

2-1-1 02-1-2 02-1-3 02-1-4 02-1-5 02-1-6 02-1-7 02-1-8 02-1-9 02-1-10 02-1-11 02-1-12 02-1-13 02-1-14 02-1-15 0

2-1-16 02-1-17 02-1-18 02-1-19 02-1-20 02-1-21 02-1-22 02-1-23 02-1-24 02-1-25 02-1-26 02-1-27 02-1-28 02-1-29 02-1-30 02-1-31 02-1-32 0

F 2-1-33/34 0F 2-1-35/36 0F 2-1-37/38 0F 2-1-39/40 0F 2-1-41/42 0

2-1-43 02-1-44 02-1-45 02-1-46 02-1-47 02-1-48 02-1-49 02-1-50 02-1-51 02-1-52 02-1-53 02-1-54 02-1-55 02-1-56 02-1-57 02-1-58 02-1-59 02-1-60 02-1-61 02-1-62 02-1-63 02-1-64 02-1-65 02-1-66 02-1-67 02-1-68 02-1-69 02-1-70 02-1-71 0




Section 2.1 (cont)Electronic Display System

2-1-72 02-1-73 02-1-74 02-1-75 02-1-76 02-1-77 02-1-78 02-1-79 02-1-80 02-1-81 02-1-82 02-1-83 02-1-84 02-1-85 02-1-86 02-1-87 02-1-88 02-1-89 02-1-90 02-1-91 02-1-92 02-1-93 02-1-94 02-1-95 02-1-96 02-1-97 02-1-98 02-1-99 02-1-100 02-1-101 02-1-102 02-1-103 02-1-104 02-1-105 02-1-106 02-1-107 02-1-108 02-1-109 02-1-110 02-1-111 02-1-112 02-1-113 02-1-114 02-1-115 02-1-116 02-1-117 02-1-118 02-1-119 02-1-120 0

2-1-121 02-1-122 02-1-123 02-1-124 02-1-125 02-1-126 02-1-127 02-1-128 02-1-129 02-1-130 02-1-131 02-1-132 02-1-133 02-1-134 02-1-135 02-1-136 02-1-137 02-1-138 02-1-139 02-1-140 02-1-141 02-1-142 02-1-143 02-1-144 02-1-145 02-1-146 02-1-147 02-1-148 02-1-149 02-1-150 02-1-151 02-1-152 02-1-153 02-1-154 02-1-155 02-1-156 02-1-157 02-1-158 02-1-159 02-1-160 02-1-161 02-1-162 02-1-163 02-1-164 02-1-165 02-1-166 02-1-167 02-1-168 02-1-169 02-1-170 02-1-171 0




Section 2.1 (cont)Electronic Display System

2-1-172 02-1-173 02-1-174 02-1-175 02-1-176 02-1-177 02-1-178 0

F 2-1-179/180 0F 2-1-181/182 0

2-1-183 02-1-184 02-1-185 02-1-186 02-1-187 02-1-188 02-1-189 02-1-190 02-1-191 02-1-192 02-1-193 02-1-194 02-1-195 02-1-196 02-1-197 02-1-198 02-1-199 02-1-200 02-1-201 02-1-202 02-1-203 02-1-204 02-1-205 02-1-206 02-1-207 02-1-208 02-1-209 02-1-210 02-1-211 02-1-212 02-1-213 02-1-214 02-1-215 02-1-216 02-1-217 02-1-218 0

F 2-1-219/220 0F 2-1-221/222 0F 2-1-223/224 0F 2-1-225/226 0

Table of Contents - Section 2.2TC2-2-1 0TC2-2-2 0

Section 2.2AHZ-800 Attitude Heading Reference System

2-2-1 02-2-2 02-2-3 02-2-4 02-2-5 02-2-6 02-2-7 02-2-8 02-2-9 02-2-10 02-2-11 02-2-12 02-2-13 02-2-14 02-2-15 02-2-16 02-2-17 02-2-18 02-2-19 02-2-20 02-2-21 02-2-22 02-2-23 02-2-24 0


Section 2.3ADZ-850 Micro Air Data System

2-3-1 02-3-2 02-3-3 02-3-4 02-3-5 02-3-6 02-3-7 02-3-8 02-3-9 02-3-10 02-3-11 02-3-12 02-3-13 02-3-14 0




Section 2.3 (cont)ADZ-850 Micro Air Data System

2-3-15 02-3-16 0


Section 2.4AA-300 Radio Altimeter System

2-4-1 02-4-2 02-4-3 02-4-4 02-4-5 02-4-6 02-4-7 02-4-8 0


Section 2.5Weather Radar System

2-5-1 02-5-2 02-5-3 02-5-4 02-5-5 02-5-6 02-5-7 02-5-8 02-5-9 02-5-10 0

F 2-5-11/12 0F 2-5-13/14 0

2-5-15 02-5-16 02-5-17 02-5-18 02-5-19 02-5-20 02-5-21 02-5-22 02-5-23 02-5-24 02-5-25 02-5-26 02-5-27 0

2-5-28 0

Table of Contents - Section 2.6TC2-6-1 0TC2-6-2 0TC2-6-3 0TC2-6-4 0

Section 2.6SRZ-850 Integrated Radio System

2-6-1 02-6-2 02-6-3 02-6-4 02-6-5 02-6-6 02-6-7 02-6-8 02-6-9 02-6-10 02-6-11 02-6-12 02-6-13 02-6-14 02-6-15 02-6-16 02-6-17 02-6-18 02-6-19 02-6-20 02-6-21 02-6-22 02-6-23 02-6-24 02-6-25 02-6-26 02-6-27 02-6-28 02-6-29 02-6-30 02-6-31 02-6-32 02-6-33 02-6-34 02-6-35 02-6-36 02-6-37 02-6-38 02-6-39 02-6-40 0

F 2-6-41/42 0




Section 2.6 (cont)SRZ-850 Integrated Radio System

2-6-43 02-6-44 02-6-45 02-6-46 0

F 2-6-47/48 0F 2-6-49/50 0F 2-6-51/52 0

2-6-53 02-6-54 0

F 2-6-55/56 02-6-57 02-6-58 0

F 2-6-59/60 02-6-61 02-6-62 02-6-63 02-6-64 0


Section 2.7Traffic Alert and Collision AvoidanceSystem (TCAS)

2-7-1 02-7-2 02-7-3 02-7-4 02-7-5 02-7-6 02-7-7 02-7-8 02-7-9 02-7-10 02-7-11 02-7-12 02-7-13 02-7-14 0

F 2-7-15/16 02-7-17 02-7-18 02-7-19 02-7-20 02-7-21 02-7-22 02-7-23 02-7-24 0


Section 2.8Flight Management System (FMS)

2-8-1 02-8-2 02-8-3 02-8-4 02-8-5 02-8-6 02-8-7 02-8-8 02-8-9 02-8-10 02-8-11 02-8-12 02-8-13 02-8-14 02-8-15 02-8-16 0

F 2-8-17/18 02-8-19 02-8-20 02-8-21 02-8-22 0


Section 2.9Global Positioning System (GPS)

2-9-1 02-9-2 02-9-3 02-9-4 02-9-5 02-9-6 02-9-7 02-9-8 02-9-9 02-9-10 02-9-11 02-9-12 02-9-13 02-9-14 0





Section 2.10Flight Director System (FDS)

2-10-1 02-10-2 02-10-3 02-10-4 02-10-5 02-10-6 02-10-7 02-10-8 02-10-9 02-10-10 02-10-11 02-10-12 02-10-13 02-10-14 02-10-15 02-10-16 02-10-17 02-10-18 02-10-19 02-10-20 02-10-21 02-10-22 0

F 2-10-23/24 0F 2-10-25/26 0

2-10-27 02-10-28 02-10-29 02-10-30 02-10-31 02-10-32 02-10-33 02-10-34 02-10-35 02-10-36 02-10-37 02-10-38 02-10-39 02-10-40 02-10-41 02-10-42 02-10-43 02-10-44 02-10-45 0

2-10-46 02-10-47 02-10-48 02-10-49 02-10-50 02-10-51 02-10-52 02-10-53 02-10-54 02-10-55 02-10-56 02-10-57 02-10-58 02-10-59 02-10-60 02-10-61 02-10-62 02-10-63 02-10-64 0

F 2-10-65/66 0F 2-10-67/68 0

2-10-69 02-10-70 02-10-71 02-10-72 02-10-73 02-10-74 02-10-75 02-10-76 02-10-77 02-10-78 02-10-79 02-10-80 02-10-81 02-10-82 02-10-83 02-10-84 02-10-85 02-10-86 02-10-87 02-10-88 02-10-89 02-10-90 02-10-91 02-10-92 02-10-93 02-10-94 02-10-95 02-10-96 02-10-97 02-10-98 0




Section 2.10 (cont)Flight Director System (FDS)

2-10-99 02-10-100 02-10-101 02-10-102 02-10-103 02-10-104 0


Section 2.11Autopilot/Yaw Damper System

2-11-1 02-11-2 02-11-3 02-11-4 02-11-5 02-11-6 02-11-7 02-11-8 02-11-9 02-11-10 02-11-11 02-11-12 02-11-13 02-11-14 02-11-15 02-11-16 02-11-17 02-11-18 0

F 2-11-19/20 0F 2-11-21/22 0

2-11-23 02-11-24 0

F 2-11-25/26 02-11-27 02-11-28 02-11-29 02-11-30 0

F 2-11-31/32 02-11-33 02-11-34 02-11-35 02-11-36 0

F 2-11-37/38 02-11-39 0

2-11-40 0F 2-11-41/42 0

2-11-43 02-11-44 02-11-45 02-11-46 0


TABLE OF CONTENTS

SECTION/PARAGRAPH/TITLE VOLUME I PAGE


- INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRO-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRO-1

2. Reference Documents . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRO-1

3. How This Manual Is Organized . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRO-2

4. Critical Items Compliance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRO-3

5. Abbreviations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRO-4

1 SYSTEM OVERVIEW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-1

2. System Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-13

A. Electronic Display System (EDS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-14

B. AHZ-800 Attitude Heading Reference System (AHRS) . . . . . . . . . . . . 1-15

C. ADZ-850 Micro Air Data System . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-16

D. AA-300 Radio Altimeter System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-16

E. PRIMUS® 650/660 Weather Radar System . . . . . . . . . . . . . . . . . . . . . 1-16

F. PRIMUS® II Integrated Radio System . . . . . . . . . . . . . . . . . . . . . . . . 1-17

G. Traffic Alert and Collision Avoidance System (TCAS) . . . . . . . . . . . . 1-17

H. Flight Management System (FMS) - Optional . . . . . . . . . . . . . . . . . . . 1-18

I. Global Positioning System (GPS) - Optional . . . . . . . . . . . . . . . . . . . 1-18

J. Flight Director System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-19

K. Autopilot/Yaw Damper System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-20

3. Digital Data Buses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-21

A. Radio System Bus (RSB) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-21

B. Digital Audio Bus . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-29

C. Commercial Standard Digital Bus (CSDB) . . . . . . . . . . . . . . . . . . . . . 1-31

D. ARINC 429 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-31

(1) Field Definitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-31

(2) Label - Bits 1 thru 8 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-32

(3) Data Field - Bits 11 thru 29 . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-32

(4) Sign Status Matrix (Bits 30 and 31) . . . . . . . . . . . . . . . . . . . . 1-34

(5) Parity (Bit 32) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-34

(6) Waveform Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-35

E. RS-422 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-36

F. RS-232 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-36

G. Serial Control Interface (SCI) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-37

H. Weather Radar Picture Data (WXPD) . . . . . . . . . . . . . . . . . . . . . . . . . 1-37

I. Integrated Computer Bus (ICB) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-37

J. SG/DU Bus . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-37

Page TC-122-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.

TABLE OF CONTENTS (Cont)



2 SYSTEM DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1

2.1 ELECTRONIC DISPLAY SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-1

2. Component Descriptions and Locations . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-4

A. IC-600 Integrated Avionics Computer . . . . . . . . . . . . . . . . . . . . . . . 2-1-4

B. DU-870 Display Units . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-7

(1) Video and Dimming System . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-8

(2) System Monitor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-8

C. BL-870 Bezel Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-9

(1) IN/HPA Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-10

(2) STD (Standard) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-10

(3) BARO Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-10

D. BL-871 Bezel Controller (-831 and -851) . . . . . . . . . . . . . . . . . . . . . 2-1-11

E. DC-550 Display Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-13

(1) FULL/WX Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-15

(2) GSPD/TTG Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-15

(3) ET Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-15

(4) NAV Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-15

(5) FMS Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-15

(6) Bearing (BRG) Source Select Knobs . . . . . . . . . . . . . . . . . . 2-1-16

(7) Radio Altitude (RA) Set Knob . . . . . . . . . . . . . . . . . . . . . . . 2-1-17

(8) System TEST Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-17

F. GC-550 Guidance Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-18

(1) Flight Director 1 (FD1) Button . . . . . . . . . . . . . . . . . . . . . . 2-1-20

(2) Course 1 (CRS 1) Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-20

(3) Heading (HDG) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-21

(4) Heading (HDG) Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-21

(5) Navigation (NAV) Button . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-21

(6) Approach (APR) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-21

(7) Bank (BNK) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-22

(8) Autopilot (AP) Engage Button . . . . . . . . . . . . . . . . . . . . . . 2-1-22

(9) Couple (CPL) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-22

(10) Yaw Damper (YD) Button . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-22

(11) Speed (SPD) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-23

(12) Vertical Speed (VS) Button . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-23

(13) Speed (SPD) Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-23

(14) Flight Level Control (FLC) Button . . . . . . . . . . . . . . . . . . . . 2-1-24





2.1 ELECTRONIC DISPLAY SYSTEM (cont)

2. Component Descriptions and Locations (cont)

F. GC-550 Guidance Control Unit (cont)

(15) Altitude (ALT) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-24

(16) Altitude Select (ASEL) Knob . . . . . . . . . . . . . . . . . . . . . . . . 2-1-24


(18) Course 2 (CRS 2) Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-25

G. Reversionary Panels (Embraer) . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-26

H. DA-800 Data Acquisition Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-27

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-29

A. Electronic Display System Bus Interface . . . . . . . . . . . . . . . . . . . . 2-1-29

B. PFD/MFD Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-30

(1) Bezel Controller Interface . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-30

(2) Guidance Control Unit Interface . . . . . . . . . . . . . . . . . . . . . 2-1-31

(3) Master Warning/Caution Light Interface . . . . . . . . . . . . . . . 2-1-31

C. PFD Attitude Director Indicator Operation . . . . . . . . . . . . . . . . . . . 2-1-43

(1) ADI Sphere . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-43

(2) Attitude Source Annunciations . . . . . . . . . . . . . . . . . . . . . . 2-1-47

(3) Autopilot Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-48

(4) Flight Director Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-49

(5) Vertical Deviation Display . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-55

(6) Marker Beacons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-57

(7) Radio Altitude Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-58

(8) Radio Altitude Minimums Display . . . . . . . . . . . . . . . . . . . . 2-1-58

(9) Excessive Attitude Declutter . . . . . . . . . . . . . . . . . . . . . . . 2-1-59

D. PFD Horizontal Situation Indicator Operation . . . . . . . . . . . . . . . . 2-1-60

(1) Heading Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-60

(2) Lateral Deviation Display . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-65

(3) To/From Pointer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-68

(4) Course Select/Desired Track Display . . . . . . . . . . . . . . . . . 2-1-69

(5) Drift Bug . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-70

(6) Bearing Pointers and Source Identifiers . . . . . . . . . . . . . . . 2-1-70

(7) Distance Readout . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-72

(8) Distance Identifier (FMS) . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-73

(9) Time-To-Go Readout . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-73

(10) Ground Speed Readout . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-73

(11) Elapsed Time Readout . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-73

(12) Wind Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-74

(13) Weather Radar Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-74






3. Operation (cont)

E. PFD Air Data Display Operation . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-79

(1) Altitude Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-79

(2) Airspeed Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-82

(3) Vertical Speed Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-86

F. PFD TCAS Display Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-89

(1) TCAS Vertical Speed Indications (VSI) . . . . . . . . . . . . . . . . 2-1-89

(2) TCAS Mode Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-92

G. PFD Miscellaneous Annunciators . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-93

(1) Air Data Computer Source Annunciator . . . . . . . . . . . . . . . 2-1-93

(2) Air Data Computer Test Annunciator . . . . . . . . . . . . . . . . . 2-1-93

(3) AHRS Test Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-93

(4) Symbol Generator Annunciator . . . . . . . . . . . . . . . . . . . . . 2-1-96

(5) Navigation Source Annunciator . . . . . . . . . . . . . . . . . . . . . 2-1-96

(6) FMS Cross-Track Mode Annunciator . . . . . . . . . . . . . . . . . 2-1-96

(7) FMS Accuracy Annunciator . . . . . . . . . . . . . . . . . . . . . . . . 2-1-97

(8) FMS Status Annunciators . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-98

(9) FMS Message Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-98

(10) Vertical Track Alert Annunciator . . . . . . . . . . . . . . . . . . . . 2-1-98

(11) ILS Approach Category Annunciator . . . . . . . . . . . . . . . . . 2-1-98

(12) Windshear Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-99

H. PFD Comparison Monitor Annunciators . . . . . . . . . . . . . . . . . . . 2-1-100

(1) Indicated Airspeed (IAS) Comparison Monitor Annunciator 2-1-100

(2) Pitch (PIT) and Roll (ROL) Attitude Comparison Monitor

Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-100

(3) Radio Altitude (RA) Comparison Monitor Annunciator . . . 2-1-100

(4) Glideslope (GS) and Localizer (LOC) Comparison Monitor

Annunciators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-100

(5) CAS Comparison Monitor Annunciator . . . . . . . . . . . . . . . 2-1-102

(6) Heading (HDG) Comparison Monitor Annunciator . . . . . . . 2-1-102

(7) Altitude (ALT) Comparison Monitor Annunciator . . . . . . . 2-1-102

I. PFD Test Mode Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-102

(1) Weight-On-Wheels Test Mode Display Formats . . . . . . . . . 2-1-102

(2) Not Weight-On-Wheels Test Mode Display Formats . . . . . 2-1-104






3. Operation (cont)

J. Multifunction Display Operation . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-105

(1) General Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-106

(2) MFD Bezel Menus . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-108

(3) Joystick . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-122

(4) Common MFD Map/Plan Format Data . . . . . . . . . . . . . . . . 2-1-123

(5) FMS Map/Plan Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-130

(6) MFD Map Format Display . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-133

(7) MFD Plan Format Display . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-138

(8) MFD TCAS Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-140

(9) MFD Checklist Display . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-144

(10) System Page Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-154

(11) MFD Test Mode Display . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-175

K. Engine Instrument Crew Alerting System (EICAS) Display Interface 2-1-177

L. EICAS Display Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-183

(1) Engine Instrument Section . . . . . . . . . . . . . . . . . . . . . . . . 2-1-187

(2) Cabin and APU Section . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-198

(3) Crew Alerting System Message Section . . . . . . . . . . . . . . 2-1-201

(4) Flight Control Information Section . . . . . . . . . . . . . . . . . . 2-1-209

M. EICAS Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-213

N. Reversionary Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-215

(1) Symbol Generator Reversion . . . . . . . . . . . . . . . . . . . . . . 2-1-215

(2) Sensor Reversion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-215

(3) Display Unit Reversion . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-217





2.2 AHZ-800 ATTITUDE HEADING REFERENCE SYSTEM . . . . . . . . . . . . . . . . . . . . . . . 2-2-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-1


A. AH-800 Attitude Heading Reference Unit (AHRU) . . . . . . . . . . . . . . . 2-2-2

B. Memory Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-4

C. FX-600 Flux Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-5

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-6

A. Pilot’s AHRS Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-6

B. Copilot’s AHRS Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-6

C. Modes of Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-10

(1) Initialization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-10

(2) Full Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-10

(3) DG Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-10

(4) Basic Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-11

(5) Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-11

(6) Maintenance Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-12

D. Reversionary Switching . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-12

4. Fault Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-13

A. Fault Detection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-13

B. Power-On BITE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-14

C. Continuous BITE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-15

D. Fault Reaction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-16

(1) Critical . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-16

(2) Non-Critical . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-16

E. Maintenance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-16

F. Fault Storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-16

G. Flight Faults . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-16

(1) Fault Service Status . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-17

(2) Miscellaneous Status . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-17

(3) Fault Type . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-17

H. Ground Faults . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-17

I. Fault Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-18

(1) PFD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-18

(2) MFD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-20

(3) EICAS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-23





2.3 ADZ-850 MICRO AIR DATA SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-1


A. AZ-850 Micro Air Data Computer (MADC) . . . . . . . . . . . . . . . . . . . . 2-3-2

B. BL-870 PFD Bezel Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-4

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-5

A. Pilot’s Air Data System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-5

B. Copilot’s Air Data System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-7

C. Modes of Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-9

D. MADC Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-9

E. Static Source Error Correction (SSEC) . . . . . . . . . . . . . . . . . . . . . . 2-3-9

F. Operational Range . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-10

G. Overspeed Warning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-10


A. Primary Flight Display (PFD) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-13

B. Multifunction Display (MFD) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-15

C. EICAS Display (EICAS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-16

2.4 AA-300 RADIO ALTIMETER SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-1


3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-4

A. Auxiliary Radio Altitude Output . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-4

B. Primary Radio Altitude Output . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-4

C. RA Minimum Annunciation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-4

D. Low Altitude Awareness . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-5

E. Radio Altitude Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-5

4. Fault Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-8





2.5 WEATHER RADAR SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-1


A. Weather Radar Receiver Transmitter Antenna (RTA) . . . . . . . . . . . . 2-5-2

B. Weather Radar Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-6

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-9

A. Target Alert (TGT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-9

B. Rain Echo Attenuation Compensation Technique (RCT) Mode . . . . 2-5-15

C. Turbulence (TRB) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-15

D. Test (TST) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-15

E. Weather (WX) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-15

F. Ground Map (GMAP) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-15

G. Standby (SBY) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-16


A. PFD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-17

B. MFD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-18

C. PRIMUS® 650/870 Weather Radar Test Mode . . . . . . . . . . . . . . . . . 2-5-19

D. PRIMUS® 660/880 Weather Radar Test Mode . . . . . . . . . . . . . . . . . 2-5-21

(1) On-Ground TEST Display (with TEXT FAULTS Enabled) . . . 2-5-21

(2) In Flight TEST Display (with TEXT FAULTS Enabled) . . . . . 2-5-22

(3) Fault Monitors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-22

2.6 SRZ-850 INTEGRATED RADIO SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-1


A. RM-855 Radio Management Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-2

(1) RMU Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-6

(2) Backup Navigation Display . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-8

(3) Backup Engine Display . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-20

B. CD-850 Clearance Delivery Control Head (Tuning Backup Control

Head) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-23

(1) System Installation Annunciator . . . . . . . . . . . . . . . . . . . . . 2-6-24

(2) Remote Tune Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-24

(3) Tuning Cursor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-24

(4) NAV AUDIO On Annunciator . . . . . . . . . . . . . . . . . . . . . . . . 2-6-24

(5) Emergency (EMRG) Mode Annunciator . . . . . . . . . . . . . . . . 2-6-24

(6) Squelch (SQ) Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25





2.6 SRZ-850 INTEGRATED RADIO SYSTEM (cont)

2. Component Descriptions and Locations (cont)


Head) (cont)

(7) Transmit (TX) Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25

(8) NAV AUDIO On/Off Switch . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25

(9) Squelch (SQ) On/Off Switch . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25

(10) Tuning Knobs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25

(11) Normal/Emergency Mode Switch . . . . . . . . . . . . . . . . . . . . 2-6-25

(12) Transfer Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25

(13) Radio Tuning Annunciators . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25

C. AV-850A Audio Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-26

(1) COM1, COM2, COM3, and HF Microphone Switches . . . . . . 2-6-27

(2) Passenger Address (PAX) Microphone Switch . . . . . . . . . . 2-6-27

(3) Emergency (EMER) Switch . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-27

(4) Audio Source Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-27

(5) ID/Voice Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-28

(6) Speaker and Headphone Controls . . . . . . . . . . . . . . . . . . . 2-6-28

(7) Sidetone (ST) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-28

(8) Marker (MKR) Beacon Volume Control . . . . . . . . . . . . . . . . 2-6-28

(9) Marker Beacon MUTE and HI/LO SENS Control . . . . . . . . . . 2-6-28

(10) Interphone (INPH) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-28

D. RCZ-851(x) Integrated Communications Unit . . . . . . . . . . . . . . . . . 2-6-29

E. RNZ-851(x) Integrated Navigation Unit . . . . . . . . . . . . . . . . . . . . . . 2-6-32

F. AT-860 ADF Combined Sense/Loop Antenna . . . . . . . . . . . . . . . . . 2-6-35

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-36

A. RM-855 Radio Management Unit Interface Diagram . . . . . . . . . . . . 2-6-38

B. CD-850 Clearance Delivery Head Interface . . . . . . . . . . . . . . . . . . . 2-6-43

C. AV-850A Audio Control Unit Interface . . . . . . . . . . . . . . . . . . . . . . 2-6-44

D. Communications Unit Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-53

E. Navigation Unit Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-57


A. PFD Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-61

B. RMU Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-61





2.7 TRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM (TCAS II) . . . . . . . . . . . . 2-7-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-1


A. RT-910 TCAS Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-2

B. AT-910 Directional Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-4

C. Typical Bottom Omnidirectional Antenna . . . . . . . . . . . . . . . . . . . . 2-7-5

D. Other Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-5

E. TCAS/MFD Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-6

F. TCAS/RMU Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-6

(1) Line Select Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-6

(2) CODE Line Select Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-6

(3) MODE Line Select Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-7

(4) RANGE Line Select Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-8

(5) SURVEILLANCE WINDOW Line Select Key . . . . . . . . . . . . . . 2-7-8

(6) PGE Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-8

G. TCAS Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-10

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-11

A. TCAS Computer Unit Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-12

B. TCAS ARINC 429 Output Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-13

C. TCAS Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-17

(1) Traffic Advisories . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-17

(2) Resolution Advisories . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-20

D. TCAS Display Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-22

(1) Traffic Advisory (TA) Logic . . . . . . . . . . . . . . . . . . . . . . . . 2-7-22

(2) Resolution Advisory (RA) Logic . . . . . . . . . . . . . . . . . . . . . 2-7-22

E. TCAS Preflight Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-22

(1) Activate TCAS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-22

(2) Activate TCAS Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-22

(3) Perform Self-test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-23



B. Fault Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-24





2.8 FLIGHT MANAGEMENT SYSTEM (FMS) - Optional . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-1


A. NZ-2000 Navigation Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-2

B. IM-803 Configuration Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-4

C. CD-810 Control Display Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-5

D. DL-900 Data Loader . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-10

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-11

A. FMS Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-13

B. FMS ARINC 429 Input/Output Data . . . . . . . . . . . . . . . . . . . . . . . . 2-8-14



B. MFD Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-21

2.9 GLOBAL POSITIONING SYSTEM (GPS) - Optional . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-1


A. Global Positioning System Sensor Unit . . . . . . . . . . . . . . . . . . . . . . 2-9-2

B. CD-810 Control Display Unit (CDU) GPS Status . . . . . . . . . . . . . . . . 2-9-3

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-6

A. Modes of Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-6

(1) Self-Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-6

(2) Initialization Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-6

(3) Acquisition Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-6

(4) Navigation (NAV) Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-7

(5) Aided Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-7

(6) Altitude Aiding Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-8

(7) Fault Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-8

(8) Mode Provisioning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-8

B. GPS Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-8

C. ARINC 429 Input Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-10

D. ARINC 429 Output Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-11






2.10 FLIGHT DIRECTOR SYSTEM (FDS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-1

A. Flight Director Data Management . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-2

B. Flight Director Couple Switching . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-2

C. Master/Slave Air Data Target Switching . . . . . . . . . . . . . . . . . . . . . 2-10-2

D. Flight Director Mode Synchronization . . . . . . . . . . . . . . . . . . . . . . 2-10-3

E. Flight Director Mode Annunciation . . . . . . . . . . . . . . . . . . . . . . . . 2-10-3

F. Flight Director Command Bar Logic . . . . . . . . . . . . . . . . . . . . . . . 2-10-3

G. Altitude Preselect Function . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-3

2. Component Descriptions and Locations . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-4

A. IC-600 Integrated Avionics Computer . . . . . . . . . . . . . . . . . . . . . . 2-10-4

B. GC-550 Guidance Panel Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-6

(1) Flight Director Mode Switches . . . . . . . . . . . . . . . . . . . . . . 2-10-8

(2) Heading (HDG) Select Knob . . . . . . . . . . . . . . . . . . . . . . . . 2-10-8

(3) Course (CRS) Select Knob . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-9

(4) Altitude Select Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-9

(5) Couple (CPL) Pushbutton . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-9

(6) Bank Pushbutton (BNK) . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-9

(7) FD1/FD2 Pushbuttons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-9

C. DC-550 Display Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-10

D. PC-400 Autopilot Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-12

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-14

A. Flight Director Functions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-14

(1) PFD Command Bars . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-14

(2) GS CAP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-15

(3) GS Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-15

(4) Lateral Beam Sensor (LBS) . . . . . . . . . . . . . . . . . . . . . . . 2-10-15

(5) LOC/BC CAP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-15

(6) LOC/BC TRACK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-15

(7) True Airspeed (TAS) Gain Programming . . . . . . . . . . . . . . 2-10-16

(8) Vertical Beam Sensor (VBS) . . . . . . . . . . . . . . . . . . . . . . . 2-10-16

(9) Vertical Path Gain Programming . . . . . . . . . . . . . . . . . . . 2-10-16

(10) VOR CAP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-16

(11) VOR Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-16

(12) VOR OSS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-17

(13) VOR AOSS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-17





2.10 FLIGHT DIRECTOR SYSTEM (FDS) (cont)

3. Operation (cont)

B. Flight Director Lateral (Roll) Channel Functional Operation . . . . . 2-10-17

(1) Flight Director Lateral (Roll) Modes Interface . . . . . . . . . . 2-10-17

(2) Heading Select (HDG) Mode . . . . . . . . . . . . . . . . . . . . . . . 2-10-27

(3) Heading Select Mode Engage/Reset/Disengage Logic . . . . 2-10-28

(4) VOR (NAV) Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-29

(5) VOR Approach (VAPP) Mode . . . . . . . . . . . . . . . . . . . . . . 2-10-36

(6) VOR/VAPP Engage/Reset/Disengage Logic . . . . . . . . . . . . 2-10-37

(7) Localizer (NAV) and Back Course (BC) Modes . . . . . . . . . 2-10-41

(8) Localizer/Back Course Mode Engage/Reset/Disengage

Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-51

(9) Long Range Navigation . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-54

(10) LNAV Mode Engage/Reset/Disengage Logic . . . . . . . . . . . 2-10-57

C. Flight Director Vertical (Pitch) Channel Functional Operation . . . 2-10-59

(1) Flight Director Vertical (Pitch) Modes Interface . . . . . . . . 2-10-59

(2) Pitch Attitude Hold . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-69

(3) Pitch Attitude Hold Mode Engage/Reset/Disengage Logic . 2-10-70

(4) Vertical Speed (VS) Hold Mode . . . . . . . . . . . . . . . . . . . . . 2-10-71

(5) Vertical Speed (VS) Hold Mode Engage/Reset/Disengage

Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-73

(6) Speed (SPD) Select Mode . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-74

(7) Speed (SPD) Select Mode Engage/Reset/Disengage Logic 2-10-76

(8) Flight Level Change (FLC, FLCH) Mode . . . . . . . . . . . . . . 2-10-77

(9) Flight Level Change (FLC) Mode Engage/Reset/Disengage

Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-79

(10) Altitude Preselect (ASEL) Mode . . . . . . . . . . . . . . . . . . . . 2-10-80

(11) Altitude Preselect (ASEL) Mode Engage/Reset/Disengage

Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-87

(12) Altitude Hold (ALT) Mode . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-88

(13) Altitude Hold (ALT) Mode Engage/Reset/Disengage Logic 2-10-90

(14) ILS Approach (APR) Mode . . . . . . . . . . . . . . . . . . . . . . . . 2-10-91

(15) ILS Approach (APR) Mode Engage/Reset/Disengage Logic 2-10-97

(16) Go-Around (GA) Mode (Wings Level) . . . . . . . . . . . . . . . . 2-10-99

(17) Go-Around (GA) Mode Engage/Reset/Disengage Logic . . . 2-10-101

(18) Windshear Mode (WSHR) . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-102

(19) Windshear (WSHR) Mode Engage/Reset/Disengage Logic . 2-10-103





2.11 AUTOPILOT/YAW DAMPER SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-1

A. Autopilot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-1

B. Yaw Damper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-2

C. AP/YD System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-2

2. Component Descriptions and Locations . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-4


B. GC-550 Guidance Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-7

(1) Autopilot Engage Pushbutton . . . . . . . . . . . . . . . . . . . . . . . 2-11-9

(2) Yaw Damper Engage Pushbutton . . . . . . . . . . . . . . . . . . . . 2-11-9

(3) CPL Pushbutton . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-9

C. PC-400 Autopilot Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-10

D. SM-200 Servo Drive and SB-201 Servo Bracket . . . . . . . . . . . . . . 2-11-12

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-15

A. Modes of Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-15

(1) Heading Hold and Wings Level . . . . . . . . . . . . . . . . . . . . . 2-11-15

(2) Roll Hold . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-15


(4) Flight Director Couple and Lift Compensation . . . . . . . . . 2-11-16

(5) Turn Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-17

(6) Pitch Wheel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-17

(7) Touch Control Steering (TCS) . . . . . . . . . . . . . . . . . . . . . . 2-11-17

B. Autopilot/Yaw Damper Engage Logic . . . . . . . . . . . . . . . . . . . . . . 2-11-23

(1) Yaw Damper Engagement . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-23

(2) Autopilot Engagement . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-24

C. Roll Axis Autopilot Servo Loop . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-27

(1) SM-200 Roll Servo Drive and Bracket . . . . . . . . . . . . . . . . 2-11-27

(2) IC-600 Integrated Avionics Computer (IAC) . . . . . . . . . . . . 2-11-28

D. Pitch Axis Autopilot Servo Loop . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-33

(1) SM-200 Elevator Servo Drive and Bracket . . . . . . . . . . . . . 2-11-33


E. Pitch Axis Autopilot Trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-36

F. Yaw Damper Rudder Axis Servo Loop . . . . . . . . . . . . . . . . . . . . . 2-11-39


(2) SM-200 Servo Drive and Bracket . . . . . . . . . . . . . . . . . . . 2-11-40





2.11 AUTOPILOT/YAW DAMPER SYSTEM (cont)

4. Fault Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-43

A. Autopilot/Yaw Damper Monitoring Overview . . . . . . . . . . . . . . . . 2-11-43

B. Hardover Malfunction Protection . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-43

C. System Response to Failures . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-43

D. Monitor Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-44

(1) Pitch Servo Position Monitor . . . . . . . . . . . . . . . . . . . . . . 2-11-44

(2) Primary Pitch Attitude Comparison . . . . . . . . . . . . . . . . . 2-11-44

(3) Secondary Pitch Attitude Comparision Monitor . . . . . . . . 2-11-44

(4) Normal Accelertion Monitor . . . . . . . . . . . . . . . . . . . . . . . 2-11-44

(5) Roll Servo Position Monitor . . . . . . . . . . . . . . . . . . . . . . . 2-11-44

(6) Primary Roll Attitude Comparison Monitor . . . . . . . . . . . . 2-11-45

(7) Secondary Roll Attitude Comparison Monitor . . . . . . . . . . 2-11-45

(8) Roll Rate Monitor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-45

(9) Yaw Servo Position Monitor . . . . . . . . . . . . . . . . . . . . . . . 2-11-45

(10) Auto Trim Runaway Monitor . . . . . . . . . . . . . . . . . . . . . . . 2-11-45

(11) Auto Trim Inoperative Monitor . . . . . . . . . . . . . . . . . . . . . 2-11-45

(12) Autopilot/Yaw Damper Disconnect Monitor . . . . . . . . . . . . 2-11-46



SECTION/PARAGRAPH/TITLE VOLUME II PAGE


3 SYSTEM INTERCONNECTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-1

4 MAINTENANCE PRACTICES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-1

2. Equipment and Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-3

3. Procedure for the AH-800 Attitude Heading Reference Unit (AHRU) . . . . . . . . 4-4

4. Procedure for the AHRU Mounting Tray Fan Filter . . . . . . . . . . . . . . . . . . . . . 4-5

5. Procedure for the AHRU Mounting Tray Fan . . . . . . . . . . . . . . . . . . . . . . . . . . 4-6

6. Procedure for the AHRU Mounting Tray . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-8

7. Procedure for the AT-860 ADF Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-9

8. Procedure for the AT-910 TCAS Directional Antenna . . . . . . . . . . . . . . . . . . 4-10

9. Procedure for the AV-850A Audio Control Unit . . . . . . . . . . . . . . . . . . . . . . . 4-12

10. Procedure for the AZ-850 Micro Air Data Computer (MADC) . . . . . . . . . . . . . 4-19

11. Procedure for the BL-870/871 Bezel Assembly . . . . . . . . . . . . . . . . . . . . . . . 4-20

12. Procedure for the CD-810 Control Display Unit . . . . . . . . . . . . . . . . . . . . . . 4-21

13. Procedure for the CD-850 Clearance Delivery Control Head (CDH) . . . . . . . . 4-21

14. Procedure for the DA-800 Data Acquisition Unit (DAU) . . . . . . . . . . . . . . . . . 4-23

15. Procedure for the DC-550 Display Controller . . . . . . . . . . . . . . . . . . . . . . . . 4-24

16. Procedure for the DL-900 Data Loader . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-25

17. Procedure for the DU-870 Display Units . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-26

18. Procedure for the FX-600 Thin Flux Valve . . . . . . . . . . . . . . . . . . . . . . . . . . 4-28

19. Procedure for the GC-550 Guidance Control Unit . . . . . . . . . . . . . . . . . . . . . 4-43

20. Procedure for the Global Navigation System Sensor Unit (GNSSU) . . . . . . . 4-44

21. Procedure for the IC-600 Integrated Avionics Computer (IAC) . . . . . . . . . . . 4-45

22. Procedure for the IM-803 Installation Module . . . . . . . . . . . . . . . . . . . . . . . . 4-47

23. Procedure for the MM-260 AHRS Memory Module . . . . . . . . . . . . . . . . . . . . . 4-49

24. Procedure for the NZ-2000 FMS Navigation Computer . . . . . . . . . . . . . . . . . 4-50

25. Procedure for the PC-400 Autopilot Controller . . . . . . . . . . . . . . . . . . . . . . . 4-54

26. Procedure for the RCZ-851(X) Integrated Communications Unit . . . . . . . . . . 4-55

27. Procedure for the RM-855 Radio Management Unit (RMU) . . . . . . . . . . . . . . 4-56

28. Procedure for the RNZ-851(X) Integrated Navigation Unit . . . . . . . . . . . . . . . 4-58

29. Procedure for the RNZ-851/RCZ-851 Strap Board Assembly . . . . . . . . . . . . . 4-62

30. Procedure for the RT-300 Radio Altimeter Receiver Transmitter . . . . . . . . . . 4-63

31. Procedure for the RT-910 TCAS Computer Unit . . . . . . . . . . . . . . . . . . . . . . 4-65

32. Procedure for the SM-200 Servo Drive and SB-201 Drum and Bracket

Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-66

33. Procedure for the WC-6X0/8X0 Weather Radar Controller . . . . . . . . . . . . . . . 4-68

34. Procedure for the WU-6X0/8X0 Antenna and Receiver Transmitter Unit

(RTA) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-69



SECTION/PARAGRAPH/TITLE VOLUME II PAGE


5 SHIPPING, HANDLING AND STORAGE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-1

6 HONEYWELL SUPPORT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-1

1. Worldwide Exchange/Rental Program for Regional Airlines . . . . . . . . . . . . . . 6-1

2. Contracted Maintenance Agreements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-1

3. Test Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-1

4. Customer Service Order Desk (End Item LRUs) . . . . . . . . . . . . . . . . . . . . . . . 6-2

5. Customer Support - Material (Repair Piece Part Spares Services) . . . . . . . . . 6-2

6. Warranty . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-3

7. Customer Engineering . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-3

8. Training . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-3

9. Honeywell Facilities . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-4

10. Regional Airline Customer Engineering Locations . . . . . . . . . . . . . . . . . . . . . 6-6

VOLUME III

7 SYSTEM TEST AND FAULT ISOLATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-1



List of Illustrations

FIGURE/TITLE VOLUME I PAGE


Figure 1-1. PRIMUS® 1000 System Flow Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-5/6

Figure 1-2. Cockpit Layout . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-7/8

Figure 1-3. Instrument Panel and Pedestal Component Locations . . . . . . . . . . . . . . . . . 1-9/10

Figure 1-4. PRIMUS® 1000 Component Locations . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-11/12

Figure 1-5. Radio System Bus . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-22

Figure 1-6. RSB Data Field Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-24

Figure 1-7. Digital Audio Data Sequence . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-30

Figure 1-8. Octal Label 274 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-32

Figure 1-9. Data Field (Bits 11 thru 29) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-32

Figure 1-10. BCD Bit Assignments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-33

Figure 1-11. Selected Course Data Word . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-33

Figure 1-12. DME Distance Data Word . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-33

Figure 1-13. Present Position Longitude Data Word . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-34

Figure 1-14. ARINC 429 Transmission Waveforms . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-35

Figure 2-1-1. IC-600 Integrated Avionics Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-4

Figure 2-1-2. DU-870 Display Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-7

Figure 2-1-3. BL-870 Bezel Controller (-921) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-9

Figure 2-1-4. BL-871 Bezel Controllers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-11

Figure 2-1-5. DC-550 Display Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-13

Figure 2-1-6. GC-550 Flight Guidance Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-18

Figure 2-1-7. Reversionary Panels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-26

Figure 2-1-8. DA-800 Data Acquisition Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-27

Figure 2-1-9. Electronic Display System Interface Diagram . . . . . . . . . . . . . . . . . . . . . . 2-1-33

Figure 2-1-10. Pilot’s PFD/MFD Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-34

Figure 2-1-11. Copilot’s PFD/MFD Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-35

Figure 2-1-12 (Sheet 1). Guidance Control Unit Interface Diagram . . . . . . . . . . . . . . . . . 2-1-36


Figure 2-1-13 (Sheet 1). Primary Flight Display - ADI Display Formats . . . . . . . . . . . . . . 2-1-44


Figure 2-1-14. Pitch Attitude Scale . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-46

Figure 2-1-15. Pitch Limit Indicator Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-53

Figure 2-1-16. Primary Flight Display - HSI Display Formats (Full Heading

Compass) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-61

Figure 2-1-17. Primary Flight Display - HSI Display Formats (Partial

Heading Compass) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-62



List of Illustrations (cont)



Figure 2-1-18. Partial Heading Compass Display With Weather Radar

Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-76

Figure 2-1-19. PFD Altitude Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-80

Figure 2-1-20. PFD Airspeed Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-83

Figure 2-1-21. PFD Vertical Speed Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-87

Figure 2-1-22. PFD TCAS Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-90

Figure 2-1-23 (Sheet 1). PFD Miscellaneous Annunciations . . . . . . . . . . . . . . . . . . . . . . 2-1-94


Figure 2-1-24. Comparison Monitor Annunciators . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-101

Figure 2-1-25. PFD Familiarization Test Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-103

Figure 2-1-26. BL-871 MFD Bezel Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-105

Figure 2-1-27. MFD Bezel Menu Tree . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-106

Figure 2-1-28. MFD INOP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-107

Figure 2-1-29. MFD Main Menu Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-108

Figure 2-1-30. System Page Menu Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-111

Figure 2-1-31. MFD Menu Display with FMS Installed . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-114

Figure 2-1-32. MFD Menu Display without FMS Installed . . . . . . . . . . . . . . . . . . . . . . . 2-1-114

Figure 2-1-33. Joystick Menu Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-116

Figure 2-1-34. Vspeeds Menu Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-118

Figure 2-1-35. Checklist Menu Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-120

Figure 2-1-36. Common MFD Map/Plan Format Data . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-124

Figure 2-1-37. Map Symbols . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-131

Figure 2-1-38. MFD Map Format Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-134

Figure 2-1-39. MFD Plan Format Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-139

Figure 2-1-40. MFD TCAS Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-141

Figure 2-1-41. MFD Checklist Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-145

Figure 2-1-42. Disclaimer Page . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-149

Figure 2-1-43. Normal Procedures Index Page . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-150

Figure 2-1-44. Waypoint Listing Page . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-151

Figure 2-1-45. Normal Checklist Page . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-152

Figure 2-1-46. Emergency Procedures Index Page . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-153

Figure 2-1-47. Electrical System Page Format - Normal Conditions . . . . . . . . . . . . . . . 2-1-155

Figure 2-1-48. Electrical System Page - Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-158

Figure 2-1-49. Hydraulic System Page Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-159

Figure 2-1-50. Hydraulic System Page - Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-161

Figure 2-1-51. Takeoff System Page Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-162

Figure 2-1-52. Takeoff System Page - Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-166






Figure 2-1-53. Environmental Control System Page Format . . . . . . . . . . . . . . . . . . . . . 2-1-167

Figure 2-1-54. Environmental Control System Page - Test Mode . . . . . . . . . . . . . . . . . 2-1-170

Figure 2-1-55. Fuel System Page Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-171

Figure 2-1-56. Fuel System Page - Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-174

Figure 2-1-57. MFD Test Mode Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-176

Figure 2-1-58 (Sheet 1). EICAS Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-179


Figure 2-1-59 (Sheet 1). EICAS Display Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-184



Figure 2-1-60. ITT Arc Default . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-190

Figure 2-1-61. ITT Arc During Engine Start . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-191

Figure 2-1-62. ITT Arc With Engine Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-191

Figure 2-1-63. EICAS Familiarization Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-214

Figure 2-1-64. Symbol Generator Reversion Mode Interface Diagram . . . . . . . . . . . . . 2-1-219

Figure 2-1-65. MADC Reversion Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-220

Figure 2-1-66. AHRS Reversion Mode Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . 2-1-221

Figure 2-1-67. DAU Reversion Mode Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . 2-1-222

Figure 2-2-1. AH-800 Attitude Heading Reference Unit . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-2

Figure 2-2-2. Memory Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-4

Figure 2-2-3. FX-600 Flux Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-5

Figure 2-2-4. Pilot’s AHRS Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-7

Figure 2-2-5. Copilot’s AHRS Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-8

Figure 2-2-6. PFD AHRS Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-19

Figure 2-2-7. MFD MAP MODE AHRS Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . 2-2-21

Figure 2-2-8. MFD PLAN MODE AHRS Failure Indications . . . . . . . . . . . . . . . . . . . . . . . 2-2-22

Figure 2-3-1. AZ-850 Micro Air Data Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-2

Figure 2-3-2. BL-870 PFD Bezel Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-4

Figure 2-3-3. Pilot’s MADC Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-6

Figure 2-3-4. Copilot’s MADC Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-8

Figure 2-3-5. Vmo/Mmo Curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-10

Figure 2-3-6. PFD MADC Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-14

Figure 2-3-7. MFD MAP MODE MADC Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . 2-3-15






Figure 2-4-1. RT-300 Radio Altimeter Receiver Transmitter . . . . . . . . . . . . . . . . . . . . . . . 2-4-2

Figure 2-4-2. Single AA-300 Radio Altimeter Interface Diagram . . . . . . . . . . . . . . . . . . . . 2-4-6

Figure 2-4-3. Optional Dual AA-300 Radio Altimeter Interface Diagram . . . . . . . . . . . . . . 2-4-7

Figure 2-4-4. PFD Radio Altitude Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-8

Figure 2-5-1. Typical Weather Radar Receiver Transmitter Antenna . . . . . . . . . . . . . . . . 2-5-2

Figure 2-5-2. Maximum Permissible Exposure Level Boundary . . . . . . . . . . . . . . . . . . . . 2-5-5

Figure 2-5-3. Optional Weather Radar Controllers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-6

Figure 2-5-4. PRIMUS® 650/870 Weather Radar System Interface . . . . . . . . . . . . . . . 2-5-11/12


Figure 2-5-6. PFD Weather Radar Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-17

Figure 2-5-7. MFD Weather Radar Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-18

Figure 2-5-8. PRIMUS® 650/870 MFD Weather Radar Test Mode

Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-19

Figure 2-5-9. PRIMUS® 660/880 MFD Weather Radar Test Mode

Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-23

Figure 2-6-1. Typical RM-855 Radio Management Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-2

Figure 2-6-2. Backup Navigation Display Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-10

Figure 2-6-3. Typical TO Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-11

Figure 2-6-4. Typical FROM Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-12

Figure 2-6-5. Typical ILS Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-13

Figure 2-6-6. 90-Degree Intercept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-14

Figure 2-6-7. VOR Information Not Available . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-15

Figure 2-6-8. ILS Information Not Available . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-16

Figure 2-6-9. ADF Information Not Available . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-17

Figure 2-6-10. RSB Information Not Available . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-18

Figure 2-6-11. Heading Information Not Available . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-19

Figure 2-6-12. Backup Engine Page No. 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-21


Figure 2-6-14. Typical CD-850 Clearance Delivery Control Head . . . . . . . . . . . . . . . . . . 2-6-23

Figure 2-6-15. AV-850A Audio Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-26

Figure 2-6-16. RCZ-851(x) Integrated Communications Unit . . . . . . . . . . . . . . . . . . . . . . 2-6-29

Figure 2-6-17. RNZ-851(x) Integrated Navigation Unit . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-32

Figure 2-6-18. AT-860 ADF Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-35

Figure 2-6-19. Radio System Data Buses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-36

Figure 2-6-20. Radio Management Unit Interface Diagram . . . . . . . . . . . . . . . . . . . . . 2-6-41/42






Figure 2-6-21. CD-850 Clearance Delivery Control Head Interface Diagram . . . . . . . . . . 2-6-43

Figure 2-6-22 (Sheet 1). Audio Control Unit Interface Diagram . . . . . . . . . . . . . . . . . 2-6-47/48



Figure 2-6-23. Communication Unit Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . 2-6-55/56

Figure 2-6-24. Navigation Unit Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-59/60

Figure 2-6-25. PFD Radio System Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-62

Figure 2-6-26. RMU Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-63

Figure 2-7-1. RT-910 TCAS Computer Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-2

Figure 2-7-2. AT-910 Directional Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-4

Figure 2-7-3. Typical Bottom Omnidirectional Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-5

Figure 2-7-4. Typical RM-855 Radio Management Unit (RMU) . . . . . . . . . . . . . . . . . . . . . . 2-7-7

Figure 2-7-5. TCAS Computer Unit Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . 2-7-15/16

Figure 2-7-6. TCAS MFD Symbology (Sheet 1) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-18


Figure 2-7-8. TCAS PFD Symbology . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-21

Figure 2-8-1. NZ-2000 Navigation Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-2

Figure 2-8-2. IM-803 Configuration Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-4

Figure 2-8-3. CD-810 Control Display Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-5

Figure 2-8-4. DL-900 Data Loader (Access Door Open) . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-10

Figure 2-8-5. FMS Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-17/18

Figure 2-8-6. PFD FMS Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-20

Figure 2-8-7. MFD FMS Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-22

Figure 2-9-1. Global Positioning System Sensor Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-2

Figure 2-9-2. GPS STATUS Page 1/2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-4


Figure 2-9-4. Global Positioning System Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-9






Figure 2-10-1. IC-600 Integrated Avionics Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-4

Figure 2-10-2. GC-550 Guidance Control Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-6

Figure 2-10-3. DC-550 Display Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-10

Figure 2-10-4. PC-400 Autopilot Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-12

Figure 2-10-5. Flight Director Lateral Modes Interface - Pilot’s Side . . . . . . . . . . . . 2-10-23/24

Figure 2-10-6. Flight Director Lateral Modes Interface - Copilot’s Side . . . . . . . . . . 2-10-25/26

Figure 2-10-7. VOR ARM Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-29

Figure 2-10-8. VOR (NAV) Mode Armed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-30

Figure 2-10-9. VOR Capture Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-31

Figure 2-10-10. VOR (NAV) Mode Capture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-32

Figure 2-10-11. VOR Course Cut Limiting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-33

Figure 2-10-12. VOR Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-34

Figure 2-10-13. VOR Overstation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-35

Figure 2-10-14. Localizer ARM Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-41

Figure 2-10-15. Localizer (NAV) Mode ARM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-42

Figure 2-10-16. Localizer Capture Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-43

Figure 2-10-17. Localizer Capture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-45

Figure 2-10-18. Localizer Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-47

Figure 2-10-19. Back Course Mode Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-47

Figure 2-10-20. Back Course Arm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-48

Figure 2-10-21. Back Course Capture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-49

Figure 2-10-22. Back Course Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-50

Figure 2-10-23. Long Range Navigation Capture Pictorial and Tracking . . . . . . . . . . . . 2-10-55

Figure 2-10-24. Long Range Navigation Tracking . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-56

Figure 2-10-25. Flight Director Vertical Modes Interface - Pilot’s Side . . . . . . . . . . . 2-10-65/66

Figure 2-10-26. Flight Director Vertical Modes Interface - Copilot’s Side . . . . . . . . . 2-10-67/68

Figure 2-10-27. Vertical Speed (VS) Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-72

Figure 2-10-28. Speed Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-75

Figure 2-10-29. FLC Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-78

Figure 2-10-30. Altitude Preselect Mode Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-82

Figure 2-10-31. Prior to Descent - Altitude Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-83

Figure 2-10-32. During Descent - ASEL Armed Mode . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-84

Figure 2-10-33. Start of Flare - ASEL Capture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-85

Figure 2-10-34. Level at New Altitude - Altitude Hold Mode . . . . . . . . . . . . . . . . . . . . . 2-10-86

Figure 2-10-35. Altitude Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-89






Figure 2-10-36. ILS Approach Arm Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-92

Figure 2-10-37. ILS Approach Arm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-93

Figure 2-10-38. ILS Approach Capture Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-94

Figure 2-10-39. ILS Approach (LOC) Capture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-95

Figure 2-10-40. ILS Approach (APR) Mode Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-96

Figure 2-10-41. Go-Around Mode (Wings Level) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-100

Figure 2-10-42. Windshear Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-103


Figure 2-11-2. GC-550 Guidance Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-7


Figure 2-11-4. SM-200 Servo Drive and SB-201 Servo Bracket . . . . . . . . . . . . . . . . . . . 2-11-12

Figure 2-11-5. Pilot’s Autopilot/Yaw Damper Interface . . . . . . . . . . . . . . . . . . . . . . . 2-11-19/20

Figure 2-11-6. Copilot’s Autopilot/Yaw Damper Interface . . . . . . . . . . . . . . . . . . . . . 2-11-21/22

Figure 2-11-7. Autopilot/Yaw Damper Engage Logic . . . . . . . . . . . . . . . . . . . . . . . . 2-11-25/26

Figure 2-11-8. Autopilot Roll Axis Servo Loop . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-31/32

Figure 2-11-9. Pitch Autopilot Servo Loop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-37/38

Figure 2-11-10. Yaw Damper Servo Loop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-41/42

VOLUME II

Figure 4-1. Fan Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-7

Figure 4-2. AV-850A Audio Control Unit Adjustment Locations . . . . . . . . . . . . . . . . . . . . . 4-18

Figure 4-3. Nomograph . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-32

Figure 4-4. Checklist Loading Page . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-46

Figure 4-5. COM Unit Adjustment Locations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-56

Figure 4-6. NAV Unit Adjustment Locations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-59

Figure 4-7. RT-300 Zero Adjustment Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-64



List of Tables

TABLE/TITLE VOLUME I PAGE


Table 1-1. System Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-2

Table 1-2. Optional System Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-4

Table 1-3. RSB Message Numbers (Normal Mode) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-25/26

Table 1-4. Sign Status Matrix Bit Assignments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-34

Table 1-5. Transmission Waveform Voltages . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-36

Table 2-1-1. IC-600 Integrated Avionics Computer Leading Particulars . . . . . . . . . . . . . . 2-1-5

Table 2-1-2. DU-870 Display Unit Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-7

Table 2-1-3. BL-870 Bezel Controller Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . 2-1-9

Table 2-1-4. BL-871 Bezel Controller Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . 2-1-12

Table 2-1-5. DC-550 Display Controller Leading Particulars . . . . . . . . . . . . . . . . . . . . . . 2-1-13

Table 2-1-6. GC-550 Guidance Control Unit Leading Particulars . . . . . . . . . . . . . . . . . . 2-1-18

Table 2-1-7. DA-800 Data Acquisition Unit Leading Particulars . . . . . . . . . . . . . . . . . . . 2-1-27

Table 2-1-10. Attitude Source Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-47

Table 2-1-11. Autopilot Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-48

Table 2-1-12. Yaw Damper Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-49

Table 2-1-13. Lateral Flight Director Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-51

Table 2-1-14. Vertical Flight Director Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-51

Table 2-1-15. Priority Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-52

Table 2-1-16. Vertical Deviation Pointer Display Colors . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-55

Table 2-1-17. Glideslope Deviation Scale . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-56

Table 2-1-18. FMS Vertical Deviation Scale (GPS Valid) . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-56

Table 2-1-19. FMS Vertical Deviation Scale (GPS Invalid) . . . . . . . . . . . . . . . . . . . . . . . 2-1-57

Table 2-1-20. Heading Source Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-64

Table 2-1-21. VOR Lateral Deviation Scale . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-65

Table 2-1-22. Localizer Deviation Scale . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-66

Table 2-1-23. FMS Lateral Deviation Scale (GPS Valid) . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-67

Table 2-1-24. FMS Lateral Deviation Scale (GPS Invalid) . . . . . . . . . . . . . . . . . . . . . . . . 2-1-67

Table 2-1-25. VOR To/From Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-68

Table 2-1-26. FMS To/From Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-68

Table 2-1-27. Bearing Source No. 1 Identifier . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-71


Table 2-1-29. Weather Radar Ranges . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-75

Table 2-1-30. Color Codes for Weather Radar Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-75

Table 2-1-31. PFD WX Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-77

Table 2-1-32. WX Warning Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-78



List of Tables (Cont)



Table 2-1-33. Barometric Correction Range and Resolution . . . . . . . . . . . . . . . . . . . . . . 2-1-81

Table 2-1-34. Resolution Advisory Matrix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-91

Table 2-1-35. PFD TCAS Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-92

Table 2-1-36. ADC Source Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-93

Table 2-1-37. Navigation Source Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-96

Table 2-1-38. FMS Status Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-98

Table 2-1-39. Windshear Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-99

Table 2-1-40. Menu Key Operations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-107

Table 2-1-41. MFD WX Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-127

Table 2-1-42. Heading Source Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-136

Table 2-1-43. WX and GMAP Mode Return Colors . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-137

Table 2-1-44. PFD TCAS Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-140

Table 2-1-45. Checklist Color Assignments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-146

Table 2-1-46. Waypoint Listing Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-150

Table 2-1-47. Takeoff Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-163

Table 2-1-48. Engine Takeoff Data Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-164

Table 2-1-49. N1 Indicator Dial Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-188

Table 2-1-50. ITT Engine Start Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-190

Table 2-1-51. Engine Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-197

Table 2-1-52. Ignition Annunciation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-198

Table 2-1-53. Cabin Differential Pressure Readout . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-199

Table 2-1-54. APU Turbine Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-200

Table 2-1-55. APU Exhaust Temperature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-201

Table 2-1-56. Message Inhibit Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-208

Table 2-1-57. Landing Gear Positions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-209

Table 2-1-58. Flap Positions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-210

Table 2-1-59. Pitch Trim Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-211

Table 2-1-60. DAU Reversion States . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-216

Table 2-1-61. DU Reversion States . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-217

Table 2-2-1. AH-800 Attitude Heading Reference Unit Leading Particulars . . . . . . . . . . . . 2-2-3

Table 2-2-2. Memory Module Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-4

Table 2-2-3. FX-600 Flux Valve Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-5

Table 2-2-4. AH-800 ARINC 429 Output Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-9

Table 2-2-5. AH-800 Full Performance Accuracy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-11

Table 2-2-6. AHRS ARINC 429 Output Test Mode Data . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-12






Table 2-3-1. AZ-850 Micro Air Data Computer Leading Particulars . . . . . . . . . . . . . . . . . . 2-3-3

Table 2-3-2. BL-870 PFD Bezel Controller Leading Particulars . . . . . . . . . . . . . . . . . . . . . 2-3-4

Table 2-3-3. AZ-850 MADC Performance Accuracy . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-11

Table 2-3-4. MADC Functional Test Outputs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-12

Table 2-4-1. RT-300 Radio Altimeter Receiver Transmitter Leading

Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-3

Table 2-5-1. Weather Radar Receiver Transmitter Antenna Leading

Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-3

Table 2-5-2. Weather Radar Controller Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . 2-5-7

Table 2-5-3. WC-XXX Control Functions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-7

Table 2-5-4. Target Alert Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-9

Table 2-5-5. PRIMUS® 650/870 Fault Codes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-20


Table 2-6-1. RM-855 Radio Management Unit Leading Particulars . . . . . . . . . . . . . . . . . . 2-6-3

Table 2-6-2. CD-850 Clearance Delivery Control Head Leading Particulars . . . . . . . . . . 2-6-23

Table 2-6-3. AV-850A Audio Control Unit Leading Particulars . . . . . . . . . . . . . . . . . . . . 2-6-26

Table 2-6-4. RCZ-851(x) Integrated Communications Unit Leading

Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-30

Table 2-6-5. RNZ-851(x) Integrated Navigation Unit Leading Particulars . . . . . . . . . . . . 2-6-33

Table 2-6-6. AT-860 ADF Antenna Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-35

Table 2-7-1. RT-910 TCAS Computer Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . 2-7-2

Table 2-7-2. AT-910 Directional Antenna Leading Particulars . . . . . . . . . . . . . . . . . . . . . 2-7-4

Table 2-7-3. RT-910 TCAS Computer ARINC 429 Output Data Table . . . . . . . . . . . . . . . . 2-7-13

Table 2-7-4. RT-910 TCAS Computer-To-Mode S Transponder Data Table . . . . . . . . . . . 2-7-14

Table 2-7-5. RCZ-851E Communications Unit-To-TCAS Computer Data

Table . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-14

Table 2-7-6. MFD/TCAS Symbology . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-17

Table 2-8-1. NZ-2000 Navigation Computer Leading Particulars . . . . . . . . . . . . . . . . . . . . 2-8-3

Table 2-8-2. IM-803 Configuration Module Leading Particulars . . . . . . . . . . . . . . . . . . . . 2-8-4

Table 2-8-3. CD-810 Control Display Unit Leading Particulars . . . . . . . . . . . . . . . . . . . . . 2-8-6

Table 2-8-4. DL-900 Data Loader Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-10

Table 2-8-5. FMS Navigation Computer ARINC 429 Output Data Table . . . . . . . . . . . . . . 2-8-14






Table 2-8-6. FMS Navigation Computer Unused ARINC 429 Output Data

Table . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-15

Table 2-8-7. IAC to Navigation Computer ARINC 429 Input Data Table . . . . . . . . . . . . . . 2-8-16

Table 2-8-8. PFD Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-19

Table 2-8-9. MFD Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-21

Table 2-9-1. Global Positioning System Sensor Unit Leading Particulars . . . . . . . . . . . . 2-9-2

Table 2-9-2. GNSSU ARINC 429 Input Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-10

Table 2-9-3. GNSSU ARINC 429 Output Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-11

Table 2-10-1. IC-600 Integrated Avionics Computer Leading Particulars . . . . . . . . . . . . 2-10-5

Table 2-10-2. GC-550 Guidance Panel Unit Leading Particulars . . . . . . . . . . . . . . . . . . . 2-10-6

Table 2-10-3. DC-550 Display Controller Leading Particulars . . . . . . . . . . . . . . . . . . . . 2-10-10

Table 2-10-4. PC-400 Autopilot Controller Leading Particulars . . . . . . . . . . . . . . . . . . 2-10-13

Table 2-10-5. Heading Select Mode Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-27

Table 2-10-6. VOR/VOR Approach Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-36

Table 2-10-7. Localizer (LOC) and Back Course (BC) Mode Operating

Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-51

Table 2-10-8. LNAV Mode Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-57

Table 2-10-9. Pitch Attitude Hold Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-69

Table 2-10-10. Vertical Speed Hold Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-72

Table 2-10-11. Speed (SPD) Hold Mode Operating Limits . . . . . . . . . . . . . . . . . . . . . . . 2-10-76

Table 2-10-12. Flight Level Change (FLC) Hold Mode Operating Limits . . . . . . . . . . . . 2-10-79

Table 2-10-13. Altitude Preselect (ASEL) Mode Operating Limits . . . . . . . . . . . . . . . . . 2-10-87

Table 2-10-14. Altitude Hold (ALT) Mode Operating Limits . . . . . . . . . . . . . . . . . . . . . 2-10-89

Table 2-10-15. ILS Approach (APR) Mode Operating Limits . . . . . . . . . . . . . . . . . . . . . 2-10-96

Table 2-11-1. IC-600 Integrated Avionics Computer (Autopilot/

Yaw Damper Function) Leading Particulars . . . . . . . . . . . . . . . . . . 2-11-5

Table 2-11-2. GC-550 Flight Guidance Control Unit Leading Particulars . . . . . . . . . . . . 2-11-7


Table 2-11-4. SM-200 Servo Drive and SB-201 Servo Bracket Leading

Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-13

Table 2-11-5. Autopilot Roll Axis Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-29

Table 2-11-6. Pitch Channel Axis Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-35




TABLE/TITLE VOLUME II PAGE


Table 3-1. Interconnect Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-3

Table 4-1. AV-850A Audio Control Unit Adjustments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-17

Table 4-2. Aircraft Alignment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-31

Table 4-3. Aircraft Alignment Example 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-33



VOLUME III

Table 7-1. Integrated Maintenance Test (IMT) Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . 7-3



(Blank Page)



INTRODUCTION

1. General

This manual provides general system maintenance instructions and theory of operation for thePRIMUS® 1000 Integrated Avionics System for the Embraer 145 aircraft. It also providesinterface information and interconnect diagrams to permit a general understanding of the overallsystem.

The purpose of this manual is to help you operate, maintain and troubleshoot the PRIMUS®1000 Integrated Avionics System to the LRU level.

Common system maintenance procedures are not presented in this manual. The bestestablished shop and flight line practices should be used.

2. Reference Documents

Additional information on subsystems installed as part of the PRIMUS® 1000 IntegratedAvionics System is available in the following publications:

Document TitleHoneywell

Publication Number

PRIMUS® 1000 Integrated Avionics Pilot’s Manual A28-1146-112

PRIMUS® 870 Weather Radar Pilot’s Manual A28-1146-56

PRIMUS® 870 Weather Radar System Description andInstallation Manual

A09-3946-01

PRIMUS® II Integrated Radio System Pilot’s Manual A28-1146-50

PRIMUS® II Integrated Radio System Operation andInstallation Manual

A15-3800-01

PRIMUS® II Integrated Radio System Event CodesPocket Guide

A04-3800-01

RCZ-850 Module Installation Instructions 62-0097-000-02

RNZ-850 Module Installation Instructions 62-0096-000-02

FMZ-Series Flight Management System Pilot’s Manual 28-1146-43

Flight Management System Pocket Guide 28-1146-58

AA-300 Radio Altimeter Operation and Installation Manual 15-3321-06

Page INTRO-122-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.


Document TitleHoneywell

Publication Number

Global Navigation System Sensor Unit (GNSSU)Installation Manual

95-8698

TCZ-910 Traffic Alert and Collision Avoidance System(TCAS II) Pilot’s Manual

A28-1146-070

TCZ-910 Traffic Alert and Collision Avoidance System(TCAS II) System Description and Installation Manual

A15-3840-001

Electronic Programmable Checklist A35-3642-002-01

3. How This Manual Is Organized

The material in this manual has been arranged so that experienced maintenance and servicepersonnel can refer directly to those sections that relate to their work, while the lessexperienced reader will find the manual a valuable introduction to the PRIMUS® 1000 IntegratedAvionics System. This manual is organized into seven (7) sections as follows.

• SECTION 1 SYSTEM OVERVIEW

The purpose of this section is to give the reader a brief overview of how the entirePRIMUS® 1000 Integrated Avionics System is organized, define terms used in the sectionsthat follow, and to serve as a guide for further study of these sections. The numerous digitaldata buses used within the PRIMUS® 1000 system are also discussed here.

• SECTION 2 SYSTEM DESCRIPTION

This section is divided into 11 subsections that provide information about Honeywellmanufacured subsystems for the PRIMUS® 1000 Integrated Avionics System. Eachsubsection describes the function and interface of the line replaceable units (LRUs) of eachsubsystem, as well as subsystem operation and cockpit displays. Block diagrams ofsubsystem interface, figures, and tables are included as an aid to understanding systemoperation. The 11 subsections are listed as follows:

– Section 2.1 Electronic Display System

– Section 2.2 Attitude Heading Reference System

– Section 2.3 Air Data System

– Section 2.4 Radio Altimeter System

– Section 2.5 Weather Radar System

– Section 2.6 Radio System

– Section 2.7 Traffic/Collision Avoidance System

– Section 2.8 Flight Management System

– Section 2.9 Global Positioning System

– Section 2.10 Flight Director System

– Section 2.11 Autopilot/Yaw Damper System



• SECTION 3 SYSTEM INTERCONNECTS (Volume II)

This section provides wiring data for the PRIMUS® 1000 Integrated Avionics System. Thisinformation is provided as an aid for troubleshooting, should a failure occur during flight orground test.

• SECTION 4 MAINTENANCE PRACTICES (Volume II)

This section describes the procedures to remove and reinstall the Honeywell LRUs.Procedures are also provided to replace lamps, set screws, and knobs. Where applicable,adjustment data is also provided. This section is divided into paragraphs that are inalphabetical order according to the unit type designator.

• SECTION 5 SHIPPING, STORAGE, AND HANDLING (Volume II)

Information on shipping, storage, and handling of all system components is contained inmanual, Honeywell Pub. No. 09-1100-01.

• SECTION 6 HONEYWELL SUPPORT (Volume II)

This section briefly describes Honeywell’s worldwide exchange/rental program (commonlyreferred to as SPEX) for spares. Phone numbers and addresses of Honeywell’s supportcenters are also included for your convenience.

• SECTION 7 SYSTEM TEST AND FAULT ISOLATION (Volume II)

This section contains the ground maintenance test procedures. Use these procedures tocheck the components of the PRIMUS® 1000 Integrated Avionics System for correctinstallation and proper operation, as well as return to service test.

4. Critical Items Compliance

NOTICECRITICAL ITEMS

COMPLIANCE REQUIRED

Honeywell has an Airworthiness Analysis procedure performed for all its airborne equipment tomake sure that equipment will not cause a dangerous in-flight condition. As a result of theanalysis, specific critical parts, some steps of assembly, and some tests are identified asinstallation critical. Complete agreement with these procedures and tests is necessary to getthe approved results, certain installations have been designated INSTALLATION CRITICAL, and100 percent compliance with those installations is required.

The clearance between the keeper pins and the drum brackets, and the diameter of the aircraftcontrol cables are designated INSTALLATION CRITICAL.

Measuring the distance between the keeper pins and the servo drum bracket for properclearance, and verifying the diameter of the aircraft control cables are critical to avoid failuresthat could cause a dangerous flight condition. Specific methods of installation are required toensure that jamming of the cable by the keeper and drum is extremely improbable.

Refer to the REMOVAL/REINSTALLATION AND ADJUSTMENT instructions in this manual forprocedures on how to verify the keeper and drum clearance, and cable diameter.



5. Abbreviations

Weights and measurements in this manual use both U.S. and S.I. (metric) values. The lettersymbols for units of measurement and the abbreviations are the same as shown in ANSI/IEEEStd 260 and ASME Y1.1, except as identified below.

Abbreviation Equivalent

A/C Aircraft

ac, AC Alternating Current

AC FAA Advisory Circular

ADC Air Data Computer

ACFT Aircraft

ACH Auxillary Control Head

ADF Automatic Direction Finder

AHRS Attitude Heading Reference System

AHRU Attitude Heading Reference Unit

ALS Aircraft Lighting System

ALT Altitude

ANNUN Annunciator(s)

ANSI American National Standards Institute

AP Autopilot

ARINC Aircraft Radio Incorporated

ASCII American Standard Code for Information Interchange

ASIC Application Specific Integrated Circuit

AUX Auxiliary

AWG American Wire Gauge standard

ATCRBS Air Traffic Control Radar Beacon System

AZ Azimuth

Baro Barometric

BCD Binary-Coded Decimal

BNR Binary Data

BOSC Bottom of Step Climb

BRG Bearing

CCA Circuit Card Assembly

CCW Counterclockwise

CDH Clearance Delivery Head

CHAN Channel

Ckt Bkr Circuit Breaker

CLK Clock

COMM Communications

CPL Couple

CRC Cyclic Redundancy Check

CRS Selected Course




CRT Cathode Ray Tube

CTS Clear To Send

CW Clockwise

dc, DC Direct Current

DC Display Controller

DEV Deviation

DH Decision Height

DIST Distance

DLS Digital Lighting System

DME Distance Measuring Equipment

DP Differential Pressure

DTR Data Transmitter Ready

DU Display Unit

ECS Environmental Control System

EEPROM Electrically Erasable Programmable Read-Only Memory

EDS Electronic Display System

EFIS Electronic Flight Instrumentation System

EIA Electronic Industries Association

EICAS Engine Indicating Crew Alerting System

FAA United States Federal Aviation Administration

FADEC Full Authority Digital Engine Control

FB Feedback

FD Flight Director

FDS Flight Director System

FWC Fault Warning Computer

GA Go-Around

GCR Ground Clutter Reduction

GHz Gigahertz

GND Ground

GNSSU Global Navigational Systems Satellite Unit

GS Ground Speed

GS Glideslope

HDG Heading

HDLC High-level Data Link Control

hPa Pressure in Hectopascals

HPN Honeywell Part Number

Hz Hertz

I/O Input/Output

IAC Integrated Avionics Computer

IAS Indicated Airspeed




ICB Integrated Computer Bus

ID Identification

ILS Instrument Landing System

IMT Integrated Maintenance Test

inHg Pressure in Inches of Mercury

ITT Inter Turbine Temperature

kHz Kilohertz

LH Left Hand

LOC Localizer

LRN Long Range NAV

LRU Line Replaceable Unit

LSB Least Significant Bit

MAG Magnetic

MDA Minimum Descent Altitude

MFD Multifunction Display

MHz Megahertz

MKR Marker Beacon

MPEL Maximum Permissible Exposure Level

MSB Most Significant Bit

NAV Navigation

NOC NAV on Course

P/O Part Of

PB Pushbutton

PFD Primary Flight Display

PN Part Number

Pot Potentiometer/Rheostat

RT Receiver Transmitter

RAD ALT, RADALT Radio Altimeter

RAM Random-Access Memory

REACT Rain Echo Attenuation Compensation Turbulence

REL Relative

RES Resolution

Rev Reversion

RH Right Hand

RSB Resolution Significant Bits

RTA Receiver Transmitter Antenna

RX Receive

SAT Static Air Temperature

SCI Serial Communication Interface

SDI Source/Destination Identifier




Sel Select

SG Symbol Generator

SMT Surface Mount Technology

SRN Short Range NAV

SSEC Static Source Error Correction

SSM Sign Status Matrix

STAB Stabilization

STD Standard

T/R Transmit/Receive

TAS True Airspeed

TAT Total Air Temperature

TBA To Be Advised

TBD To Be Determined

TBS To Be Supplied

TCS Touch Control Steering

TOC Top of Climb

TOD Top of Descent

TSO Technical Standard Order

TTL Tuned to Localizer

TX Transmit

V ac Volts, alternating current

V dc Volts, direct current

VHF Very High Frequency (30 to 300 MHz)

VLSI Very Large Scale Integration

VOR VHF Omnidirectional Range

WOW Weight-On-Wheels

WX Weather Radar

WXPD Weather Radar Picture Data

YD Yaw Damper



(Blank Page)


TABLE OF CONTENTS - SECTION 1

SECTION/PARAGRAPH/TITLE PAGE


1 SYSTEM OVERVIEW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-1

2. System Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-13

A. Electronic Display System (EDS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-14

B. AHZ-800 Attitude Heading Reference System (AHRS) . . . . . . . . . . . . 1-15

C. ADZ-850 Micro Air Data System . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-16

D. AA-300 Radio Altimeter System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-16

E. PRIMUS® 650/660 Weather Radar System . . . . . . . . . . . . . . . . . . . . . 1-16

F. PRIMUS® II Integrated Radio System . . . . . . . . . . . . . . . . . . . . . . . . 1-17

G. Traffic Alert and Collision Avoidance System (TCAS) . . . . . . . . . . . . 1-17

H. Flight Management System (FMS) - Optional . . . . . . . . . . . . . . . . . . . 1-18

I. Global Positioning System (GPS) - Optional . . . . . . . . . . . . . . . . . . . 1-18

J. Flight Director System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-19

K. Autopilot/Yaw Damper System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-20

3. Digital Data Buses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-21

A. Radio System Bus (RSB) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-21

B. Digital Audio Bus . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-29

C. Commercial Standard Digital Bus (CSDB) . . . . . . . . . . . . . . . . . . . . . 1-31

D. ARINC 429 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-31

(1) Field Definitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-31

(2) Label - Bits 1 thru 8 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-32

(3) Data Field - Bits 11 thru 29 . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-32

(4) Sign Status Matrix (Bits 30 and 31) . . . . . . . . . . . . . . . . . . . . 1-34

(5) Parity (Bit 32) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-34

(6) Waveform Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-35

E. RS-422 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-36

F. RS-232 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-36

G. Serial Control Interface (SCI) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-37

H. Weather Radar Picture Data (WXPD) . . . . . . . . . . . . . . . . . . . . . . . . . 1-37

I. Integrated Computer Bus (ICB) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-37

J. SG/DU Bus . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-37

Page TC1-122-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.

TABLE OF CONTENTS - SECTION 1 (Cont)


FIGURE/TITLE PAGE


Figure 1-1. PRIMUS® 1000 System Flow Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-5/6

Figure 1-2. Cockpit Layout . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-7/8

Figure 1-3. Instrument Panel and Pedestal Component Locations . . . . . . . . . . . . . . . . 1-9/1-10

Figure 1-4. Primus®1000 Component Locations . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-11/1-12

Figure 1-5. Radio System Bus . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-22

Figure 1-6. RSB Data Field Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-24

Figure 1-7. Digital Audio Data Sequence . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-30

Figure 1-8. Octal Label 274 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-32

Figure 1-9. Data Field (Bits 11 thru 29) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-32

Figure 1-10. BCD Bit Assignments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-33

Figure 1-11. Selected Course Data Word . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-33

Figure 1-12. DME Distance Data Word . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-33

Figure 1-13. Present Position Longitude Data Word . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-34

Figure 1-14. ARINC 429 Transmission Waveforms . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-35

List of Tables

TABLE/TITLE PAGE

Table 1-1. System Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-2

Table 1-2. Optional System Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-4

Table 1-3. RSB Message Numbers (Normal Mode) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-25/26

Table 1-4. Sign Status Matrix Bit Assignments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-34

Table 1-5. Transmission Waveform Voltages . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-36

Page TC1-222-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.


SECTION 1SYSTEM OVERVIEW

1. General

The PRIMUS® 1000 Integrated Avionics System, as shown in Figure 1-1, consists of thefollowing subsystems:

• Electronic Display System (EDS)

• AHZ-800 Attitude Heading Reference System (AHRS)

• ADZ-850 Micro Air Data System

• AA-300 Radio Altimeter System

• PRIMUS® 650/660 Weather Radar System

• PRIMUS® 870/880 Weather Radar System - Optional

• PRIMUS® II Integrated Radio System

• TCZ-910 Traffic Alert and Collision Avoidance System (TCAS-II)

• FMZ-2000 Flight Management System (FMS) - Optional

• Global Positioning System (GPS) - Optional

• Flight Director System

• Autopilot/Yaw Damper System.

The PRIMUS® 1000 Integrated Avionics System also provides for automatic fault reporting andnon-intrusive monitoring of sensor data during on-ground maintenance.

The PRIMUS® 1000 Integrated Avionics System is a completely integrated, fail-passiveautopilot/yaw damper/flight director and display system, which has a full complement ofhorizontal and vertical flight guidance modes. These include all radio guidance modes, longrange navigation (LRN) system tracking, and air data oriented vertical modes.

Three-axis aircraft attitude stabilization and path control is provided throughout the aircraft’snormal flight regime. The automatic path mode commands (flight director) are generated byeither IC-800 Integrated Avionics Computer (IAC), which integrates the attitude and headingreference, air data and symbol generator functions into a complete aircraft control system. Thesingle autopilot/yaw damper is located in the pilot’s IC-600 IAC.

The PRIMUS® 1000 Integrated Avionics System also has provisions for I/O (input/output) anddata management with external radio navigation subsystems through digital/serial data businterfaces (radio systems bus). Additional data management activities which cross theboundaries of the functions listed above, include system monitoring, self-test, and failureannunciation. Both IC-600 Integrated Avionics Computers communicate with each other over adedicated IC bus.

Page 1-122-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.


Table 1-1 lists the components and part numbers that compose a standard system, andTable 1-2 lists optional subsystem components. Figure 1-1 provides an architectural diagramfor the entire PRIMUS® 1000 system, and Figure 1-2 illustrates the cockpit layout for theHoneywell equipment. Figure 1-3 illustrates the component locations of the instrument paneland the pedestal. Figure 1-4 illustrates the approximate component locations of the Honeywellequipment on the Embraer 145.

Table 1-1. System Components

System Component Qty Part Number

AircraftReferenceDesignator

AH-800 Attitude HeadingReference Unit

2 HG2010AC02 1/C1

FX-600 Flux Valve 2 7010133 4/C4

AZ-850 Micro Air Data Computer 2 7014700-918 9/C9

GC-550 Guidance Control Unit 1 7021170-951 11

SM-200 Servo (Aileron) 1 4006719-910 12

SB-201 Servo Bracket 1 4005842 12A

SM-200 Servo (Elevator) 1 4006719-910 13


SM-200 Servo (Rudder) 1 4006719-910 14


RT-300 Radio Altimeter ReceiverTransmitter

1 7001840-937 20

WU-650 Weather Radar ReceiverTransmitter Antenna

1 7008470-822 59

WU-660 Weather Radar ReceiverTransmitter Antenna

1 7021450-601 59

WC-650 Weather Radar Controller 1 7008471-607 61

DC-550 Display Controller 2 7016986-401,-501 115/C115

PC-400 Autopilot Controller 1 7003897-925 129

DU-870 Display Unit 5 7014300-901 130/C130131/C131

132

BL-870 Bezel (PFD withInclinometer)

2 7014331-921 N/A



Table 1-1. System Components

System Component Qty Part Number

AircraftReferenceDesignator

BL-871 Bezel (MFD with Buttons) 2 7014332-931 N/A

BL-871 Bezel (EICAS with Knob) 1 7014332-951 N/A

DA-800 Data Acquisition Unit 2 7013348-911 136/137

RCZ-851E Integrated Comm Unit(Diversity Mode S)

1 7510700-806 143

RCZ-851G Integrated Comm Unit(ATCRBS)

1 7510700-808 C143

RM-855 Radio Management Unit 2 7013270-931,-941 144/C144

AT-860 ADF Antenna 1 7510300-901 158

AV-850A Audio Panel 3 7511001-939 160/C160/E160

RNZ-851 Integrated Nav Unit(NAV/DME/ADF)

1 7510100-831 164

RNZ-851C Integrated Nav Unit(NAV only)

1 7510100-834 C164

CD-850 Clearance DeliveryControl Head (Tuning BackupControl Head)

1 7513000-801 165

IC-600 Integrated AvionicsComputer (with AP)

1 7017000-82401 190

IC-600 Integrated AvionicsComputer (without AP)

1 7017000-83401 C190

RT-910 TCAS Computer Unit 1 4066010-904 193

AT-910 TCAS Antenna 1 7514060-902 194



Table 1-2. Optional System Components

System Component Qty Part No.Aircraft

Reference Designator

RT-300 Radio AltimeterReceiver Transmitter

1 7001840-937 C20

DC-550 Display Controller(used with FMS)

2 7016986-501 115/C115

CD-810 Control Display Unit 1 7007549-901 120

NZ-2000 Navigation Computer 1 7018879-02006 121

DL-900 Data Loader 1 7016600-901 121

RCZ-851E Integrated CommUnit (Diversity Mode S)

1 7510700-806 C143

RCZ-851H Comm Unit (COMonly)

1 7510700-809 C143

Global Navigation SystemSensor Unit (GNSSU)

1 HG2021GD02 149

AT-860 ADF Antenna 1 7510300-901 C158

RNZ-851 Nav Unit(NAV/DME/ADF)

1 7510100-831 C164

IM-803 Installation Module(used with NZ-2000)

1 7014940-902 199

WU-870 Weather RadarReceiver Transmitter Antenna

1 7012640-921 59

WC-870 Weather RadarController

1 7008471-803 61

WU-880 Weather RadarReceiver Transmitter Antenna

1 7021450-801 59

WC-880 Weather RadarController

1 7008471-401 61



GC-550 GUIDANCE CONTROL UNIT

PC-400AUTOPILOT CONTROLLER

DA-800DAU

AZ-850MADCRT-300 RADIO

ALTMETER R/T

RCZ-851 INTCOM UNIT

AH-800AHRU

CD-850CLR DLYHEAD

RNZ-851 INTNAV UNIT

IC-600INTEGRATED

AVIONICSCOMPUTER

FX-600FLUX VALVE

DU-870 PFD DU-870 MFD DU-870 EICAS

WU-650RECEIVER/

TRASMITTER/ANTENNA

AV-850A AUDIO CONTROL UNIT

RM-855 RMU

RCZ-851 INTCOM UNIT

RNZ-851 INTNAV UNIT

AH-800AHRU

AZ-850MADC

DA-800DAU

WC-650 WX CONTROLLER

DU-870 MFD

RM-855 RMU


DU-870 PFD

AD-39738-R3@

DC-550DISPLAYCONTROLLER

DC-550DISPLAY

CONTROLLER

SCI

LEFTFADEC

T

T T

REMOTESWITCHES

429

RT-910TCASCOMPUTER

LEFTENGINE/

SYSTEMSDATA

RT-910DIRECTIONALANTENNA

429

HORIZONTALSTABILIZER

SYSTEM

SM-200AILERONSERVO

SM-200ELEVATOR

SERVO

SM-200RUDDERSERVO

ICB

EICASREVERSIONARYPANEL

T T

TRIGHTFADEC

429

RIGHTENGINE/

SYSTEMSDATA

FX-600FLUX VALVE

REMOTESWITCHES

SCI

WXPD


429

429

GLOBALPOSITIONING

SYSTEM(OPTIONAL)

FLIGHTMANAGEMENT

SYSTEM(OPTIONAL)

AURALWARNINGSYSTEM

RUDDERTRIM

SYSTEM

STALLPROTECTION

SYSTEM

ENVIRONMENTALCONTROLSYSTEM

GROUNDPROXIMITYWARNINGSYSTEM

HFCOM3

STSPKR HDPH

DME1 DME2 ID/VOICE

COM1 COM2 PAX

MICROPHONE EMER

ADF1NAV2

MKR

NAV1

MUTE

ADF2

INPH

BRG BRG

NAV FMSET

OFF

NAV 1

ADF

GSPDTTG

FULLWX

FMS OFF

NAV 2

ADF

FMS

RA TEST

MFD ADC

NORMEICASPFD

SG AHRS

REVERSIONARY PANEL

EICAS REV

DAU 1 DAU 2

TURNDESCEND

CLIMB

PITCH

CRS 1

PUSH SYNC PUSH SYNC

HDG SPD

PUSH IAS/M

ASEL CRS 2

PUSH SYNC

FD1 HDG

BNK

APR

NAV AP

CPL

YD

SPD FD2ALT

VS

FLC

SQ NAVAUDIO

MODES

COMNAVNAV AUDIO TX SQ PRENAV

RMT EMRG

123

429.5112.70 NAV ADF

360

S

W

E

N

CRSTO

N

33 3

ADF

VOR

DME

065

195

7.2360 OM

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

HFCOM3

STSPKR HDPH

DME1 DME2 ID/VOICE

COM1 COM2 PAX

MICROPHONE EMER

ADF1NAV2

MKR

NAV1

MUTE

ADF2

INPH

HFCOM3

STSPKR HDPH

DME1 DME2 ID/VOICE

COM1 COM2 PAX

MICROPHONE EMER

ADF1NAV2

MKR

NAV1

MUTE

ADF2

INPH

BRG BRG

NAV FMSET

OFF

NAV 1

ADF

GSPDTTG

FULLWX

FMS OFF

NAV 2

ADF

FMS

RA TEST

MFD ADC

NORMEICASPFD

SG AHRS

REVERSIONARY PANEL

REMOTEJOYSTICK

IC-600INTEGRATED

AVIONICSCOMPUTER

IN/HPA STD

BARO

10 10

10 10

20 20

350

9

1

260

280

220

200

240

260

0020

80

4500

43

5000

N

S

33

15

30

12

W

E

24

6

21

3

.410 M

359 CRS

ILS1

13.1 NM

25

VOR1

ADF2

HDG001 TGT

0

1

2

3

3

1

2TTG5MIN

1000

200 RA 29.92 IN

LOC HDG IAS GSAP YD

GS

FLAPS

SPLRS

0

CLD

LDG GEAR

CAB ALT

CAP P

CAB RATE

APU

ROLL

YAW

PITCH

UP

9

88.1 FLEX-TO

UP UP UP

OIL VIB

79 79 100100

PRESS TEMP LP HP

AA

N2

FF

FQ

99.9%

990 PPH

1000 LB

99.9%

990 PPH

1000 LB

ITT

N1

550

75.0

550

75.0ATTCS

88.1

REV REV

IGN IGNA A

END

FT

PSI

FPM

C

7000

7.5

-5000

100% 600

IN/HPA STD

BARO

10 10

10 10

20 20

350

9

1

260

280

220

200

240

260

0020

80

4500

43

5000

N

S

33

15

30

12

W

E

24

6

21

3

.410 M

359 CRS

ILS1

13.1 NM

25

VOR1

ADF2

HDG001 TGT

0

1

2

3

3

1

2TTG5MIN

1000

200 RA 29.92 IN

LOC HDG IAS GSAP YD

GS

GAIN RADAR TILTMIN MAX

OFF

STBYWX GMAP

FP0

PULLVAR

TST

PULLAUTO

RCT STAB TGT SECT

15

OFF

RSB RSB

REVERSIONARYPANEL

REVERSIONARYPANEL

RSB

N

S

33

15

30

12

W

E

24

6

21

3

25

300+15245

TASSATGSPD

FMSKDVT12.512

NMMIN

360

50 50

PLAB1

PLAB2

*PBD01LL01

KDVT

TGTTX-16

ELECM/PRNGHYD

RESETFUELECST/ORTN

STAB

ENGINE

REF TO TEMP:

REF FLX TEMP: -99

DOORS

REF A-ICE: OFF

C

OIL LVL 1 QT4 QT

-99 C

DOOROPEN

N

S

33

15

30

12

W

E

24

6

21

3

25

300+15245

TASSATGSPD

FMSKDVT12.512

NMMIN

360

50 50

PLAB1

PLAB2

*PBD01LL01

KDVT

TGTTX-16

ELECM/PRNGHYD

RESETFUELECST/ORTN

STAB

ENGINE

REF TO TEMP:

REF FLX TEMP: -99

DOORS

REF A-ICE: OFF

C

OIL LVL 1 QT4 QT

-99 C

DOOROPEN

RSB

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

MAX-TO MAX-TO

102.5 99.9N1

ITTN2

52095.0

850

250

145

PAGE 2 4 MSGS

145

195

910

96.7490

FF PPH

OIL P

OIL T

AT-860ADF ANTENNA

Figure 1-1. PRIMUS® 1000 System Flow Diagram

Page 1-5/622-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.


Figure 1-2. Cockpit Layout



PANEL LIGHTS

PTT

DISPLAYS

CHART HLDRFLOODLT

PILOT

PEDESTAL

PFD

MFD

OFF

BRTOFF

BRT

DIM

BRTDIM

BRT

DIM

BRTOFF

BRT

PTT

DISPLAYS

CHART HLDR

FLOODLT

MFD

BRT

OFF

BRT

DIM

BRT

DIM

BRT

DIM

DIM

BRT

OFF

BRT

EICAS

COPILOT

PFD

PANEL LT

NORM

PFD EICAS

MFD ADC AHRS SG

REVERSIONARY PANEL

PEDAL ADJ

FWD

AFT

PEDAL ADJ

FWD

AFT

FEETAIR

FEETAIR

AHRSDG

SLVD CCW

CWAHRS

DG

SLVD CCW

CW

BRAKEON

UP

DN

RELLOCK

DN

SELCAL

VHF 2 HF

Collins

ACT

MEM

XFR

MEMTX

SQON OFF

OFFSTO

TESTACT

COM

V

Honeywell

MICROPHONE

HFCOM1 COM2 COM3 PAX

NAV 1 NAV 2 ADF 1 ADF 2 DME 1 DME 2

EMER

BOOM MASK

ID/VOICE

SPKR HDPHINPHS.T. MKR MUTE

SPKRON

SPKROFF

LO HISENSSENS

FULLWX

GSPDTTG

ET NAV FMS

FMS

ADF

NAV 2

OFF

BRG HoneywellBRG

OFF

NAV 1

ADF

FMS RA TEST

FD1 HDG NAV

BNK

APR CPL

FLCSPD ALT FD2

CRS 1 HDG

PUSH SYNC PUSH SYNC

SPD

PUSH IAS/M

ASEL CRS 2

PUSH SYNC

AP

YD

VS

FULLWX

GSPDTTG

ET NAV FMS

FMS

ADF

NAV 2

OFF

BRG HoneywellBRG

OFF

NAV 1

ADF

FMS RA TEST

NORM

PFD EICAS

MFD ADC AHRS SG

REVERSIONARY PANEL

Honeywell

MICROPHONE

HFCOM1 COM2 COM3 PAX

NAV 1 NAV 2 ADF 1 ADF 2 DME 1 DME 2

EMER

BOOM MASK

ID/VOICE

SPKR HDPHINPHS.T. MKR MUTE

SPKRON

SPKROFF

LO HISENSSENS

WARN

CAUT

WARN

CAUT

0

2

34

5

7

1 2 00

10 1 3ALT

MB

OFF

P

U L L

TO

CAGE

406080

100

120

140160

180200

250

300

350

IAS

KNOTS

CHR

GMT LOC

ET / CHR

Mo Dy0

45 15

30

ET

DATE

GMTSET

LO

FLT NR

CHR

GMT LOC

ET / CHR

Mo Dy0

45 15

30

ET

DATE

GMTSET

LO

FLT NR

AIRCRAFT DATE

MAGNETIC COMPASS CALIBRATED

FOR ELECTRICAL EMERGENCY

FOR USE FOR USE FOR USE000045

090 225

180135 270

315

360

ARTEXELT

ON

ARM

EM

ER

GE

NC

Y U

SE

ON

LY

TEST/RESET

PRESS ONWAIT 1 SECOND

PRESS ARM

FMS

TO CONFIG

CHECK

THRUST RATING

CON CLB CRZ

ELEVDISC

AILDISC

PITCH TRIMPITCH TRIM

MAIN SYS BACKUP SYSCUTOUT CUTOUT

ROLL TRIM

YAW TRIM

LWD RWD

LEFT RIGHT

BACKUP

DN

UP

STALL PROTECTIONCUTOUT 1 CUTOUT 2

TEST

TEST

UP

0.

0.

9.

9.

2.

22

3.

33

45.

45

DOWN

.

MAX

THRUST SET

IDLE

FREE FREE

LOCKED LOCKED

MAX REV MAX REV

GUSTLOCK

GO AROUND GO AROUND

FR

I C T I O

N

LOCK

PRESSAND

PULL

PRESSAND

PULL

PULL

AND

ROTATE

EMERG/PARKBRAKE

CLOSE

OPEN

SPEED BRAKE

TURNDESCEND

CLIMB

P

I

T

C

H

HoneywellDIGITAL

ACTI VE FLT PLAN

MLF

SLC

KSLCARRIVAL

KSLCDEST

ARM ALTN

020 93.0NM

348 4.0NM

0910Z

0911Z

1/5

610

160/4920

4560

DSPLY DGRAD APRCHDR MSG

PERF NAV PREV NEXT FPL PROG DIR

CLRDEL

BRT

1 2 3

654

7 8 9

. 0

-

A B C D E F

G H I J K L

M N O P Q R

S T U V W

X Y Z

Honeywell

OFFSET

Collins

HF

R

TTEL SUP CAR

MODECHAN FREQ KHZ

V S

CHAN FREQ

CLAR+

TSTOFF

PGM PULL

MODE

PULL

100

AUTO PRESS

LAND ELEV (ft) ELV SET

+

-

DN

UPDUMP AUTO/MAN

ON MAN

DAU 1 DAU 2

EICAS REV

+

TRB GCR TGT SECT

MAXMIN

WXSBY

OFF

RCT

GAIN RADAR TILT

0 15+

-

PULLAUTO

GMAP

FP

TST

Honeywell

SQ DIM 1 2 STO

DMETSTPGEID

TUNE

COM1 NAV1

123.20 110.25

131.27

108.30

DME IPHX

MEMORY 3

AC/TCAS ADF1

1471 162.5

1 TA/RA ANT

TCAS DSPY 1TCAS DSPY 1

RANGE:

ALT: NORM

6

Honeywell

SQ DIM 1 2 STO

DMETSTPGEID

TUNE

102.5 N1 99.9

520

95.0

1850

250

145 OIL 33/64 C

OIL P

N2

ITT 490

96.7

1910

195

150

PAGE 2 4 MSGS

CLB CLB

FUEL

T/O

COM

NAV AUDIO TX SQ

NAV

EMRG

MODES

CLR DLY Honeywell

SQ NAVAUDIO

2

IN/HPA STD

BARO

10 10

10 10

20 20

350

9

1

260

280

220

200

240

260

0020

80

4500

43

5000

N

S

33

15

30

12

W

E

24

6

21

3

.410 M

359 CRS

ILS1

13.1 NM

25

VOR1

ADF2

HDG001 TGT

0

1

2

3

3

1

2TTG5MIN

1000

200 RA 29.92 IN

LOC HDG IAS GSAP YD

GSN

S

33

15

30

12

W

E

24

6

21

3

25

300+15245

TASSATGSPD

FMSKDVT12.512

NMMIN

360

50 50

PLAB1

PLAB2

*PBD01LL01

KDVT

TGTTX-16

ELECM/PRNGHYD

RESETFUELECST/ORTN

STAB

ENGINE

REF TO TEMP:

REF FLX TEMP: -99

DOORS

REF A-ICE: OFF

C

OIL LVL 1 QT4 QT

-99 C

DOOROPEN

FLAPS

SPLRS

0

CLD

LDG GEAR

CAB ALT

CAP P

CAB RATE

APU

ROLL

YAW

PITCH

UP

9

88.1 FLEX-TO

UP UP UP

OIL VIB

79 79 100100

PRESS TEMP LP HP

AA

N2

FF

FQ

99.9%

990 PPH

1000 LB

99.9%

990 PPH

1000 LB

ITT

N1

550

75.0

550

75.0ATTCS

88.1

REV REV

IGN IGNA A

END

FT

PSI

FPM

C

7000

7.5

-5000

100% 600

N

S

33

15

30

12

W

E

24

6

21

3

25

300+15245

TASSATGSPD

FMSKDVT12.512

NMMIN

360

50 50

PLAB1

PLAB2

*PBD01LL01

KDVT

TGTTX-16

ELECM/PRNGHYD

RESETFUELECST/ORTN

STAB

ENGINE

REF TO TEMP:

REF FLX TEMP: -99

DOORS

REF A-ICE: OFF

C

OIL LVL 1 QT4 QT

-99 C

DOOROPEN

IN/HPA STD

BARO

10 10

10 10

20 20

350

9

1

260

280

220

200

240

260

0020

80

4500

43

5000

N

S

33

15

30

12

W

E

24

6

21

3

.410 M

359 CRS

ILS1

13.1 NM

25

VOR1

ADF2

HDG001 TGT

0

1

2

3

3

1

2TTG5MIN

1000

200 RA 29.92 IN

LOC HDG IAS GSAP YD

GS

1. AV-850A AUDIO PANEL NO. 1

2. REVERSIONARY PANEL NO. 1

3. DU-870 DISPLAY UNIT NO. 1 (PFD)

4. DU-870 DISPLAY UNIT NO. 2 (MFD)

5. DC-550 DISPLAY CONTROLLER NO. 1

6. DU-870 DISPLAY UNIT NO. 3 (EICAS)

7. GC-550 GUIDANCE CONTROL UNIT

12. AV-850A AUDIO PANEL NO. 2

11. REVERSIONARY PANEL NO. 2

10. DU-870 DISPLAY UNIT NO. 5 (PFD)

9. DU-870 DISPLAY UNIT NO. 4 (MFD)

8. DC-550 DISPLAY CONTROLLER NO. 2

13. RM-855 RADIO MANAGEMENT UNIT NO. 1

14. EICAS REVERSIONARY PANEL

15. RM-855 RADIO MANAGEMENT UNIT NO. 2

16. CD-850 CLEARANCE DELIVERY CONTROL HEAD

17. WC-650/660 WEATHER RADAR CONTROLLER

18. PC-400 AUTOPILOT CONTROLLER

19. CD-810 CONTROL DISPLAY UNIT

5 87643

2

1

9 10

11

12

14

1513

16

17

18

19

AD-51724@

Figure 1-3. Instrument Panel and Pedestal Component Locations

Page 1-9/1-1022-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.


4. BELOW COCKPIT FLOOR COMPONENTS:

NOSE AVIONICS BAY COMPONENTS:

AH-800 ATTITUDE AND HEADING REFERENCE UNIT (2)

RNZ-851 INTEGRATED NAV UNIT (2)

NZ-2000 NAVIGATION COMPUTER

AZ-850 MICRO AIR DATA COMPUTER (2)

INSTRUMENT PANEL AND PEDESTAL MOUNTED COMPONENTS:

DU-870 DISPLAY UNIT (PFD, MFD AND EICAS)


AV-850A AUDIO PANEL (2)

DC-550 DISPLAY CONTROLLER (2)

PC-400 AUTOPILOT CONTROLLER

WC-650 WEATHER RADAR CONTROLLER

2.

3.

RADOME COMPONENTS:

RNZ-851 INTEGRATED NAV UNIT (2)

1.

WU-650/660 WEATHER RADAR RECEIVER/TRANSMITTER AND ANTENNA UNIT

BL-870 BEZEL CONTROLLER (MOUNTED ON PFD)

BL-871 BEZEL CONTROLLER (MOUNTED ON MFD)

REVERSIONARY PANEL (2)

CD-850 CLEARANCE DELIVERY CONTROL HEAD

RM-855 RADIO MANAGEMENT UNIT (2)

WC-650 WEATHER RADAR CONTROLLER

CD-810 CONTROL DISPLAY UNIT

12

4

IC-600 INTEGRATED AVIONICS COMPUTER (2) DA-800 DATA ACQUISITION UNIT NO. 1

5. FORWARD CABIN COMPONENTS:

RT-910 TRAFFIC COLLISION AND AVOIDANCE UNIT

GNSSU GLOBAL NAVIGATION SYSTEM SENSOR UNIT

6. CENTRAL FUSELAGE COMPONENTS:

RT-300 RADIO ALTIMETER RECEIVER TRANSMITTER

8. SM-200 SERVO DRIVE (AILERON)

9. SM-200 SERVO DRIVE (ELEVATOR)

10. SM-200 SERVO DRIVE (RUDDER)

7. FX-600 FLUX VALVE (2)

11. REAR BULKHEAD COMPONENTS:

DA-800 DATA ACQUISITION UNIT NO. 2

12

11

5

7

11

8, 9 6

7

AD-51725@

10

13. ADF2 ANTENNA

12. ADF1 ANTENNA

13

Figure 1-4. Primus®1000 Component Locations

Page 1-11/1-1222-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.


2. System Description

The PRIMUS® 1000 Integrated Avionics System line replaceable units, that are listed inTable 1-1, have been organized into the following subsystems:


• AHZ-800 Attitude Heading Reference System (AHRS)

• ADZ-850 Micro Air Data System

• AA-300 Radio Altimeter System

• PRIMUS® 650/660 Weather Radar System

• PRIMUS® II Integrated Radio System

• TCZ-910 Traffic Alert and Collision Avoidance System (TCAS-II)

• FMZ-2000 Flight Management System (FMS) - Optional

• Global Positioning System (GPS) - Optional

• Flight Director System

• Autopilot/Yaw Damper System.

The PRIMUS® 1000 system organization is centered around the concept of an integratedavionics computer (IAC) which performs the display and flight guidance functions normallyassociated with a symbol generator, flight director, and autopilot/yaw damper. These functionsare all co-located in the IC-600 IAC on separate circuit card assemblies (CCAs). The IC-600IAC reduces the number of aircraft LRUs by housing a number of independent functions in oneLRU. Some of these functions are managed by dedicated I/O hardware and some are managedby a micro-processor in conjunction with individual commands, switching logic, and drivecircuitry. As installed in this aircraft, only the pilot’s IC-600 IAC has an autopilot/yaw damperfunction.

During normal operation, the system displays heading, course, radio bearing, pitch and rollattitude, radio altitude, course deviation, glideslope deviation, to-from and DME indications.Lighted annunciators denote selected flight director modes. Pitch and roll flight director steeringcommands developed by the IC-600 IAC, are displayed on the primary flight display (PFD) in thecockpit. This computed steering information enables the pilot to reach and/or maintain thedesired flight path or attitude.

When the autopilot is engaged and coupled to either the pilot’s or copilot’s flight director, theaircraft is controlled with the same commands that are displayed on the PFD. When theautopilot is engaged and no flight director modes are active, the aircraft is controlled by the pilotin pitch and roll and by inserting commands through touch control steering (TCS).

Operation of a specific system component by the IC-600 IAC is dependent upon the system andother aircraft sensor data inputs. The IC-600 IAC uses software tests, in combination withbuilt-in-test (BIT) hardware to detect failures and determine I/O (input/output) signal validity.Based on the results of these tests, the IC-600 IAC determines if the system is capable ofproviding proper display, FD, AP, and YD mode control and/or mode annunciation. Systemmonitoring is active in all modes of operation.



A. Electronic Display System (EDS)

The EDS is comprised of the following LRUs:

• DU-870 Electronic Display Units

– Primary Flight Display (PFD) for each pilot

– Multifunction Display (MFD) for each pilot

– A Single Engine Indicating Crew Alerting System (EICAS) Display

• BL-870 Bezel Controller for each PFD

• BL-871 Bezel Controller for each MFD and the EICAS

• IC-600 Integrated Avionics Computer No. 1 and No. 2

• DC-550 Display Controller for each pilot

• GC-550 Guidance Control Unit

• DA-800 Data Acquisition Unit No. 1 and No. 2

• Reversionary Controllers.

The Primary Flight Display (PFD) displays pitch and roll attitude, heading, course/desiredtrack orientation, and flight path commands, as well as selected mode and sourceannunciations. The PFD also displays indicated airspeed, barometric altitude, verticalspeed, and radio altitude.

The Multifunction Display (MFD) is used to present a variety of data that includes: longrange navigation mapping, weather radar display, SAT, TAS, and TAT data, and TCASTraffic displays. Through the use of the BL-871 Bezel Controller buttons mounted on thefront of the display, the pilot can access menus that allow selection of system pages,electronic checklists, and Vspeeds, to mention a few.

The Engine Indicating Crew Alerting System (EICAS) is an integrated display that replacesthe majority of the traditional engine gauges and warning lights in the cockpit. The EICASDU is used to display the primary engine, essential subsystem, and crew alertingmessages in a single format.

The DA-800 Data Acquisition Unit (DAU) is the interface for engine data. DAU No. 1 is forthe left engine and aircraft sensors and DAU No. 2 is for the right engine and aircraftsensors.

The IC-600 IAC is the heart of the system and processes all data for display on the PFD,MFD, and EICAS. The symbol generator portion of the IC-600 IAC is the focal point forinformation flow for the EDS.

The DC-550 Display Controller provides the means for pilot control of various displayformats on the PFD, including bearing pointer select functions. The DC-550 DisplayController also provides a data acquisition function for the GC-550 Guidance Control Unitand BL-870/BL-871 bezels.



The GC-550 Guidance Control Unit allows for pilot inputs of selected heading and selectedcourse for lateral flight director modes. Pilot selection of IAS, MACH, and vertical speedtargets, as well as altitude preselect target is also through the GC-550 Guidance ControlUnit.

Three reversionary controllers are used in the system. Each crew member has anEmbraer-supplied reversionary panel. This panel allows each crew member to switch theon-side PFD or the EICAS display to the on-side MFD display unit. It also allows on-sideor cross-side selection of ADC, SG, and/or AHRS. The third controller is available to bothcrew members to select DAU reversionary modes.

The switching of navigation sensor data for display and for flight guidance is providedelectronically. All comparison monitoring of critical display data is done within the EDS.

B. AHZ-800 Attitude Heading Reference System (AHRS)

The AHRS is comprised of the following LRUs:

• AH-800 Attitude Heading Reference Unit (AHRU)

• FX-600 Flux Valve.

The AHRS is an all attitude, inertial sensor system which provides aircraft attitude,heading, rate of change, and acceleration data to the EDS, flight director, autopilot/yawdamper, weather radar antenna, and other aircraft systems.

The AHRS is a strapdown system that differs from a platform type system. A typicalplatform system has the following characteristics:

• Gimballed gyros with 2.5 to 3.0 degrees of freedom

• A spinning mass that is isolated from the airframe

• Output signals are displacement sensitive.

Characteristics of a strapdown system are:

• Does not use gimbals

• The spinning mass follows the airframe - it is not isolated

• The output signals are rate sensitive.

The AHRU uses three interferometric fiber optic gyros (IFOG), which measure the phaseshift produced between two beams of light traveling in opposite directions through opticalfiber wrapped around a core, in place of the spinning mass. These gyros are rate sensingand measure angular motion. The AHRU also uses three accelerometers to measurelinear motion.

The FX-600 Flux Valve detects the horizontal component of the earth’s magnetic field andprovides the long term magnetic reference for the AHRS. A flux valve compensator is notrequired in the system. Compass calibration data is computed by the AHRU and is storedin a memory module.



C. ADZ-850 Micro Air Data System

The ADZ-850 Micro Air Data System is comprised of the following LRUs/controls:

• AZ-850 Micro Air Data Computer (MADC)

• Baro set knob and standard (STD) pushbutton on the PFD bezel.

The AZ-850 MADC provides the IC-600 IAC with an ARINC 429 input of baro correctedaltitude, IAS, MACH, Vmo, TAS, TAT, SAT, and altitude rate. The baro set knob on thePFD bezel allows for pilot input of barometric pressure. The STD pushbutton on the PFDbezel allows for automatic barometric correction settings of either 29.92 inHg or 1013 hPa.The AZ-850 MADC is connected to the pitot/static and outside air temperature probes.

Air data parameters displayed on the PFD are:

• IAS/MACH

• Barometric Altitude

• Vertical Speed.

Air data parameters displayed on the MFD are:

• TAS

• SAT

• TAT.

D. AA-300 Radio Altimeter System

The AA-300 Radio Altimeter System is comprised of the RT-300 Radio Altitude ReceiverTransmitter and Embraer-supplied transmit/receive antenna. The AA-300 Radio Altimetersystem provides an output of absolute height above the terrain. This radio altitude isdisplayed on the PFDs. The radio altitude output is also used in determining that decisionheight has been reached and for gain programming in the flight director for ILSapproaches. Radio altitude is also used by TCAS to inhibit resolution advisories (RAs)when close to the ground.

E. PRIMUS® 650/660 Weather Radar System

The PRIMUS® 650/660 Weather Radar System is comprised of the following LRUs:

• WU-650/660 Receiver Transmitter Antenna Unit

• WC-650 Weather Radar Controller.

The PRIMUS® 650/660 Weather Radar System is an X-Band radar designed for weatherdetection, ground mapping, and analysis. Data is displayed on the PFD and the MFD.Storm intensity levels are displayed in bright colors against a deep black background.Areas of heaviest rainfall appear in magenta, next heaviest appear in red, rainfall ofmedium intensity appear in yellow and areas of weakest rainfall appear in green.



In the ground mapping mode, prominent landmarks are displayed that enable the pilot toidentify coastline, hilly and mountainous regions, as well as cities or even large structures.In GMAP mode, video levels of increasing reflectivity are displayed as black, cyan, yellow,and magenta.

A rain echo attenuation compensation technique (REACT) mode automatically increasesreceiver gain as a function of attenuation due to intervening rainfall. At the point where thereceiver can no longer detect levels less than red, a blue field is displayed indicating anout-of-calibration region. Target alert (TGT) mode is selected to indicate when level 3 (red)or greater weather is present in a sector beyond the currently displayed range.

Weather radar mode selection, range, and tilt control are provided by the WC-650 WeatherRadar Controller.

F. PRIMUS® II Integrated Radio System

The PRIMUS® II Integrated Radio System is comprised of the following LRUs:

• RM-855 Radio Management Unit

• AV-850A Audio Control Unit

• CD-850 Clearance Delivery Head (Tuning Backup Control Head)

• RCZ-851E/851G/851H Integrated Communications Unit

• RNZ-851/851C Integrated Navigation Unit.

The PRIMUS® II Integrated Radio System is a dual, remote-mounted digital radio systemwhich encompasses all standard navigation and communication functions, including VOR,DME, ILS, and VHF communications. Marker beacon and transponder (mode A/C/S)depending on installation is also included. All control functions are operated from twoRM-855 Radio Management Units (RMUs). A CD-850 Clearance Delivery Control Head isalso part of the system. Interface with the traffic alert and collision avoidance system(TCAS) is from the RMU and through the diversity Mode S transponder.

G. Traffic Alert and Collision Avoidance System (TCAS)

The TCAS is comprised of the following LRUs:

• RT-910 TCAS Computer Unit (CU)

• AT-910 TCAS Directional Antenna

• RCZ-851E Integrated COM Unit (with diversity Mode S transponder)

• RM-855 Radio Management Unit (RMU).

The TCAS CU interrogates Mode A/C/S transponders on aircraft in the vicinity and listensfor the transponder replies. By computer analysis of these replies, the airborne CUdetermines which aircraft represent potential collision threats and provides appropriatedisplay indications (or advisories) to the flight crew to ensure vertical separation. Verticalseparation is based upon predictions from own aircraft altitude data and intruder altitudedata supplied to the CU. The appropriate maneuver is one that ensures adequate verticalseparation while causing the least deviation of the TCAS aircraft from its current verticalrate.



If the threat aircraft is itself equipped with TCAS, a coordination procedure via the air-to-airMode S data link is performed before displaying the advisory to the flight crew. Thisprocedure assures that the advisories displayed in each aircraft are compatible.

Mode selection and other operational commands for the TCAS system are generated withinthe PRIMUS® II RMU, which is also part of the PRIMUS® 1000 Integrated AvionicsSystem. These commands are conveyed to the TCAS CU, from the Mode S transponder,via ARINC 429 data. If the CU calculates that an advisory should be presented to the flightcrew, it will provide ARINC 429 output data to both IACs for display, and a synthesizedvoice will be applied to the audio control units.

H. Flight Management System (FMS) - Optional

The optional FMS is comprised of the following components:

• NZ-2000 Navigation Computer (NZ)

• CD-810 Control Display Unit (CDU)

• DL-900 Data Loader.

The FMS provides lateral and vertical navigation guidance for display and coupling (lateralonly) to the AFCS. To provide high accuracy long-range navigation, the FMS computer isdesigned to connect to AHRS, GPS, and/or VOR/DME. With links to the onboardnavigation sensors, the computer develops an FMS position based on a blend or mix of thesensors. The fundamental purpose of the FMS is to provide navigation information relativeto a selected geographically located point. Navigation management allows the flight crewto define a route from the aircraft present position to any point in the world.

The system will output advisory information and steering signals to allow the flight crew orthe AFCS to steer the aircraft along a desired route. Routes are defined from the aircraftpresent position to a destination waypoint via a direct great circle route or via a series ofgreat circle legs connected by intermediate waypoints.

The CD-810 CDU is the pilot interface with the FMS. The CRT displays relative flightinformation to the pilot. The pilot enters alphanumeric data into the system via the fullalphanumeric keyboard.

The DL-900 Data Loader is used to transfer navigation data and custom data to the NZ.

I. Global Positioning System (GPS) - Optional

The optional GPS is comprised of the Global Navigation System Sensor Unit (GNSSU)component.

The GNSSU is a 12-channel receiver that receives the L1 transmissions (centered at1575.42 MHz) from the NAVSTAR GPS satellite constellation. The GPS tracks a minimumof four satellites, processes the received signals, and determines the system latitude,longitude, altitude, time, and velocity. When less than four satellites remain trackable, thesystem uses inertial information from the IRS and air data information from the MADCs tocontinue determination of position. When a fourth satellite is acquired, the system revertsto normal tracking mode.



J. Flight Director System

The flight director system is comprised of the following components:

• IC-600 Integrated Avionics Computer (IAC)

• GC-550 Guidance Controller.

The PRIMUS® 1000 Flight Director System features an integrated avionics computerconcept which combines the normal EDS display function with the flight director function.This level of integration provides a number of benefits over existing systems and greatlysimplifies the interface requirements of the flight director function. This level of integrationimplies that if the EDS is operational, the flight director will be operational. Conversely, ifthe EDS has failed, the flight director will also be failed.

Input data requirements for the flight director are totally encompassed by the EDS function.By combining the flight director and EDS processors, the flight director I/O hardware andsoftware can be virtually eliminated.

The flight director provides computed steering commands to the autopilot and for displayon the PFD. If the autopilot is not engaged, the pilot can manually fly the steeringcommand presented on the PFD. The flight director provides both lateral (roll) and vertical(pitch) steering commands. One lateral and one vertical flight director mode can be activesimultaneously. Other flight director modes can be armed to automatically become activeat the proper time.

For the flight director to do its job, it looks at the following:

• What is the pilot’s desired attitude/position/heading/etc?

• What is the aircraft’s actual attitude/position/heading/etc?

• If there is a difference between desired and actual data, correct for the difference andcontrol the speed at which the correction takes place.

The flight director computes pitch and roll steering commands based on data from a varietyof sources including:

• Air data

• Pitch and roll attitude

• Flight management system

• Magnetic heading

• VOR/DME/ILS

• Pilot inputs

• Radio altimeter.



Flight director steering commands provide a key data point in the Display and FlightGuidance System. These steering commands are output to the following subsystems:

• EDS for pilot display

• Autopilot for automatic flight path control

• Autopilot monitors.

The IC-600 IAC processes course, selected heading, attitude, air data, DME, and radionavigation data to provide computed pitch and roll steering commands for display on thePFD and for autopilot automatic flight path steering through control of the flight controlsurfaces through the SM-200 Servo Motors.

Flight director mode selection and annunciation is accomplished through mode selectbuttons on the GC-550 Guidance Controller. The flight director command cue on the PFDalso reflects the selected mode.

Flight director couple switching between the pilot’s and copilot’s flight director isaccomplished through the GC-550 Guidance Controller.

K. Autopilot/Yaw Damper System

The Autopilot/Yaw Damper System is comprised of the following LRUs:

• IC-600 Integrated Avionics Computer (IAC) - Pilots only

• GC-550 Guidance Controller

• SM-200 Servos

• Trim Power Amplifier (not Honeywell)

• PC-400 Turn/Pitch Controller.

The PRIMUS® 1000 system autopilot is a fail-passive design featuring digital attitude andservo loops. The autopilot provides aircraft stabilization and tracking of pitch and rollsteering commands from the flight director. The autopilot is not aware of which flightdirector mode(s) if any are active. The autopilot simply tracks the pitch and roll steeringcommands from the selected flight director as attitude changes.

The yaw damper provides yaw rate damping only and makes no effort to control the flightpath of the aircraft. Servo position reference is synchronized to zero at engagement and isconstantly washed out to ensure that steady state rudder forces are zero. If the ruddertrim position changes due to pilot input or aircraft configuration changes, the rudderwashes out the steady state force and allows rudder servo resynchronization.

The autopilot/yaw damper monitors are capable of disengaging the autopilot and yawdamper as an independent function. Data used in the autopilot/yaw damper computationsare processed in a manner consistent with autopilot flight-safety requirements while alsomaximizing autopilot availability. The autopilot/yaw damper engage and disengage processis also monitored to ensure that the actual engage situation at the servos correctly reflectsthe engage function status in software.



The pitch axis autopilot trim function works to maintain the aircraft attitude against longterm attitude disturbances, such as fuel burn and passenger movement.

For the autopilot to do its job, it looks at the following:

• What is the pilots desired attitude?

• What is the aircrafts actual attitude?

• If there is a difference between desired and actual data, correct for the difference andcontrol the speed at which the correction takes place.

The autopilot and yaw damper engagement is accomplished through the GC-550 GuidanceController. Basic pitch and roll movements can also be made using the PC-400 Turn/PitchController.

3. Digital Data Buses

An essential function of the PRIMUS® 1000 Integrated Avionics System is informationinterchange between subsystems and/or between LRUs within a subsystem.

Some of this information is in the form of discrete data. Discrete data is carried on a single wireand typically switches between +28 V dc and open, or between ground and open. This switcheddata is used for annunciators, warnings, and anywhere that simple "condition information" willsuffice. This is a small portion of the total information interchange.

Most of the information transfer between subsystems is accomplished through the use of digitaldata buses. The data buses found in the PRIMUS® 1000 Integrated Avionics System include:

• Radio System Bus (RSB)

• Digital Audio Bus

• Commercial Standard Digital Bus (CSDB)

• ARINC 429

• RS-422

• RS-232

• Serial Control Interface (SCI)

• Weather Radar Picture Data (WXPD)

• Integrated Computer Bus (ICB)

• SG/DU Bus.

The following paragraphs describe the operation and uses of each of the above buses.

A. Radio System Bus (RSB)

The Honeywell radio system bus (RSB), as shown in Figure 1-5, is the principalcommunications network interconnecting the LRUs in the PRIMUS® II Integrated Radiosubsystem. All the LRUs in the radio system, except the audio control units, are connectedto the RSB. Specific details regarding the operation of the radio system is coveredelsewhere in this manual.



Reliable transfers of data via RSB are ensured by designed-in redundancy and predefinedprotection and isolation mechanisms. Control and data protocols are also predefined toensure consistent application of the data bus. It is a fail-operational data bus system andactually consists of three shielded-twisted-pairs. These are the PRIMARY bus, LEFT-SIDESECONDARY bus, and RIGHT-SIDE SECONDARY bus. "Fail-operational" means that ifany device connected to the bus fails, the bus remains operational.

RMU NO. 1

120

IAC NO. 1

COM NO. 1

NAV NO. 1

CL DEL UNIT

RMU NO. 2

IAC NO. 2

COM NO. 2

NAV NO. 2

J1

M

N

J1

V

W

J1B

70

71J1C

36

37

J1

74

61

J1

10

11

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

J1A

J1B

71

81

88

102

J1

g

R

C144

C190

C143

C164

J1

V

W

M

N

36

37

70

71

10

11

71

81

88

102

74

61

J1

J1B

J1C

J1

J1

J1B

J1A

PRI RSB

PRI RSB

PRI RSB

PRI RSB

RIGHT SEC RSB

RIGHT SEC RSB

RIGHT SEC RSB

RIGHT SEC RSB

144

190

143

164

LEFT SEC RSB

PRI RSB

LEFT SEC RSB

PRI RSB

LEFT SEC RSB

PRI RSB

LEFT SEC RSB

PRI RSB

PRI RSB

165

AD-49528@

LE

FT

SE

C B

US

PR

I B

US

RIG

HT

SE

C B

US

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

120 120 120

120120

W W WW W

WW

Figure 1-5. Radio System Bus



All units that are connected to the RSB [radio management unit (RMU), integrated avionicscomputer (IAC), remote NAV unit, etc.] are defined as users. The RSB users are alltransformer coupled and impedance-matched to the data bus transmission lines. The busis a shielded-twisted-pair which is differentially driven. Data transmitted onto the busdrives one line more positive, and the other line more negative. This interface methodprovides protection from faults, transients, and RF interference. By design, the RSBinterfaces are virtually immune to lightning-induced transients, hot shorts, ground shorts,and RF threats. The design precludes any fault propagation (via RSB) between the variousinterconnected users. At the same time, the RSB interconnect structure provides superiorRF emissions characteristics, ensuring that RSB will not interfere with sensitive receiverson-board the aircraft. The users are connected to the data buses via a splicingarrangement (using solder rings) which experience has shown to be extremely reliable anddamage resistant. The type of cable that is specified for use meets regulatory guidelinesfor flammability and smoke, and is resistant to hydraulic fluids and fuel.

Data flow on RSB is bidirectional with a bit transmission rate of 2/3 MHz (1.5 µs/bit). Datatraffic flow on RSB does not require a bus controller. All users receive and identify all busdata. Since each user knows its own user number, it sets up an internal timer, based uponthe last message received, and transmits at the appropriate time. Each RSB user (otherthan those described as "Listen Only") outputs its message on the PRIMARY and its"ON-SIDE" SECONDARY buses simultaneously. This arrangement provides each user withdual-path access to its own-side data and single-path access to all cross-side data. It alsomakes it impossible for any single-point fault (such as a fire-ax or a projectile) to disable allthree data buses. For example, a failure of the PRIMARY bus will merely disablecross-side tuning of the radios, and will cause no other problems.

The clearance delivery CDH and the IC-600 are "Listen Only" devices. They DO NOTtransmit on the RSB.

Each bus user’s transmitters are safety interlocked to ensure that no user can broadcastoutside its allotted time slot or in response to another user’s request. The user interlockmechanisms effectively keep the bus users from competing for simultaneous bus timewindows, and thereby ensure reliable data flow.

A field is defined as a 192 millisecond time period that contains a sequence of24 messages spaced 8 milliseconds apart, starting with message 0 (transmits address 0)and progressing in sequence to message number 23. Thus, there are 24 possiblemessage time slots for this bus.

As shown in Figure 1-6, in the message 0 time slot, the left side NAV unit transmits onboth the PRIMARY and LEFT-SIDE SECONDARY buses. Then, in the message 1 timeslot, the right side NAV unit transmits on both the PRIMARY and RIGHT-SIDESECONDARY buses. Then, there is a spare time slot (message 2) for future expansion.Since some messages combine data from more than one radio function, RMU, COM,Transponder, VOR/LOC, Glideslope, Marker, DME, ADF, and MLS require eight messagesper system side. Left side system = 8, right side system = 8, and spare time slots = 8more, totaling 24.



/ /

/ /

/ /

/ /

0 1 2 3 4 5 18 19 20 21 22 23 0MESSAGE NO.

PRIMARY BUS

LEFT-SIDE SECONDARY

RIGHT-SIDE SECONDARY

EXPANSION TIME SLOT

NAVDATA

RMUDATA

L R S L R S L = LEFT SIDER = RIGHT SIDES = SPARE TIME SLOT

192 MSEC PER FIELD

AD-34563@

Figure 1-6. RSB Data Field Structure

When message number 23 is completed, the cycle begins again with message number 0,and the cycle repeats for as long as the system has power applied. During initialpower up, the RMUs are programmed to start the bus activity by transmitting messages 3or 4, depending on which RMU comes on line first. The sequence of transmissions isfixed, and any LRU user that is not installed in the aircraft will still have a time slotassigned at the appropriate time in the field. Therefore, removal of a unit will not disablethe bus functions.

Table 1-3 shows the message content for each message in the sequence in the normaloperational mode.

The data format of the messages on the RSB is similar to high-level data link control(HDLC). This format is described by International Standard ISO 3309-1979 (E).



Table 1-3. RSB Message Numbers (Normal Mode)

WORDPOS.

0, 1, 2Nav Rem

3, 4, 5RMU Com

6, 7, 8Nav Rem

9, 10, 11Com Rem

12, 13, 14Nav Rem

15, 16, 17RMU Nav

18, 19, 20Nav Rem

21, 22, 23IAC/FMS

1 Low MSG. NO. MSG. NO. MSG. NO. MSG. NO. MSG. NO. MSG. NO. MSG. NO. MSG. NO.1 High CONTROL CONTROL CONTROL CONTROL CONTROL CONTROL CONTROL CONTROL

2 Low MLS COM ADF COM MLS ADF ADF ATC2 High OUTPUT AZ OPMODE REL BRG STATUS OUTPUT AZ OPMODE REL BRG OPMODE

3 Low MLS COM ADF COM MLS ADF ADF ATC3 High OUTPUT GP CHAN MAG BRG CHAN OUTPUT GP CHAN MAG BRG REPLY CODE

4 Low MLS COM VOR/ILS COM MLS ADF VOR/ILS MISC.4 High AZ DEV PRESET BRG/LOC DEV PRESET AZ DEV PRESET BRG/LOC DEV STATUS

5 Low MLS ATC LEFT VOR/ILS COM MLS VOR/ILS VOR/ILS ATC5 High GP DEV OPMODE GS DEV GP DEV OPMODE GS DEV ALTITUDE

6 Low DME ATC LEFT VOR/ILS COM DME VOR/ILS VOR/ILS VHF COM6 High DIST REPLY CODE MARKER DIST CHAN MARKER OPMODE

7 Low RT-SIDE ATC RIGHT DME DIST ATC RT-SIDE VOR/ILS DME STA VHF COM7 High DME OPMODE RT-SIDE STATUS DME PRESET FMS "a" CHANNEL

8 Low DIST ATC RIGHT PRESET ATC DIST VOR-DME DME CHAN VOR/ILS8 High FMS "a" REPLY CODE DME DIST REPLY CODE FMS "a" OPMODE FMS "a" OPMODE

9 Low DME ATC/TCAS LFT-SIDE ATC DME VOR-DME DME GS VOR/ILS9 High DIST OPMODE PRESET DATA DIST CHAN FMS "a" CHANNEL

10 Low FMS "b" ATC/TCAS DME STATUS ATC FMS "b" MLS DME TTS VOR-DME10 High DME ALT/RANGE R-S PRESET ALTITUDE DME OPMODE FMS "a" OPMODE

11 Low DIST COM STRAPS DME CHAN ATC DIST MLS DME STATUS VOR-DME11 High LFT-SIDE WORD 1 R-S PRESET LFT-SIDE CHAN FMS "b" CHANNEL

12 Low DME STATUS COM STRAPS DME GS ATC/TCAS DME STATUS MLS FWD. DME CHAN MLS-DME12 High LFT-SIDE WORD 2 R-S PRESET STATUS RT-SIDE SEL. AZ FMS "b" OPMODE

13 Low DME CHAN COM STRAPS DME TTS ATC/TCAS DME CHAN MLS SEL. GP DME GS MLS-DME13 High LFT-SIDE WORD 3 R-S PRESET ALT/RANGE RT-SIDE MLS BKWD. FMS "b" CHANNEL

14 Low DME GS COM STRAPS DME IDENT AUX1 DME GS SEL. AZ DME TTS FMS "a"14 High LFT-SIDE WORD 4 R-S PRESET STATUS RT-SIDE RES FOR DME FMS "b" DME OPMODE

15 Low DME TTS DME IDENT AUX1 DME TTS MLS-DME MLS FMS "a"15 High LFT-SIDE R-S PRESET RT-SIDE OPMODE STATUS DME CHAN



Table 1-3 (cont). RSB Message Numbers (Normal Mode)

WORDPOS.

0, 1, 2Nav Rem

3, 4, 5RMU Com

6, 7, 8Nav Rem

9, 10, 11Com Rem

12, 13, 14Nav Rem

15, 16, 17RMU Nav

18, 19, 20Nav Rem

21, 22, 23IAC/FMS

16 Low DME IDENT DME STATUS AUX1 DME IDENT MLS-DME MLS FMS "b"16 High LFT-SIDE L-S PRESET RT-SIDE CHAN CHAN DME OPMODE

17 Low DME IDENT DME CHAN AUX1 DME IDENT NAV STRAPS MLS FWD. FMS "b"17 High LFT-SIDE L-S PRESET RT-SIDE WORD 1 SEL. AZ DME CHAN

18 Low VOR/ILS DME GS AUX2 MLS AUX NAV STRAPS MLS SEL. GP MLS18 High STATUS L-S PRESET STATUS DATA WORD 1 WORD 2 MLS GSTATUS OPMODE

19 Low VOR/ILS DME TTS AUX2 MLS AUX NAV STRAPS MLS BKWD. MLS19 High CHAN L-S PRESET DATA WORD 1 WORD 3 SEL. AZ. CHANNEL

20 Low VOR/ILS DME IDENT AUX2 MLS AUX NAV STRAPS MLS BASIC MLS20 High PRESET L-S PRESET DATA WORD 2 WORD 4 1,3,4,5,6 FORW/BACK

21 Low VOR/ILS DME IDENT COM CLUSTER MLS AUX AHRS-A429 MLS BASIC21 High IDENT L-S PRESET STRAPS DATA WORD 2 NAV HEADING 1,3,4,5,6 MLS GP

22 Low VOR/ILS AUX1 ADF ATC MLS AUX MLS BASIC ADF22 High IDENT OPMODE STATUS CONFIG DATA WORD 3 WORD 2 OPMODE

23 Low AUX1 ADF ATC MLS AUX MLS BASIC ADF23 High CHAN CONFIG DATA WORD 3 WORD 2 CHANNEL

24 Low NAV CLUSTER AUX1 ADF ATC MLS AUX MLS COM CLUSTER24 High STRAPS PRESET CONFIG DATA WORD 4 GEN DATA OPMODE

25 Low NAV CLUSTER AUX1 ADF COM CLUSTER MLS AUX MLS NAV CLUSTER25 High STATUS IDENT STATUS DATA WORD 4 GEN DATA OPMODE

26 Low NAV CLUSTER AUX2 ADF COM CLUSTER26 High STRAPS OPMODE IDENT STRAPS

27 Low NAV CLUSTER AUX2 COM CLUSTER SYSTEM27 High STRAPS STRAPS ON/OFF

28 Low NAV CLUSTER AUX2 COM CLUSTER POST SYS28 High STRAPS STRAPS POST RADIOS

29 Low NAV CLUSTER AUX2 ADF COM CLUSTER MLS MISC CONTRL29 High STRAPS CONFIG STRAPS CONFIG FMCS CONTR

30 Low VOR COM CLUSTER ADF COM DME NAV CLUSTER30 High CONFIG OPMODE CONFIG CONFIG CONFIG OPMODE

31 Low CHECKSUM CHECKSUM CHECKSUM CHECKSUM CHECKSUM CHECKSUM CHECKSUM CHECKSUM31 High CHECKSUM CHECKSUM CHECKSUM CHECKSUM CHECKSUM CHECKSUM CHECKSUM CHECKSUM



B. Digital Audio Bus

The PRIMUS® 1000 system uses a digital data bus to carry digital audio information fromthe remote radio system line replaceable units (LRUs) to the flight crew’s audio panels.Digitizing the audio offers the advantage of complete independence from groundingproblems within the aircraft and the absolute elimination of ground noise pick-up, whine,and cross-talk.

Each side has a "One-Way" digital audio bus, consisting of a shielded-twisted-pair which isdifferentially driven, feeding up to four audio panels. Data transmitted onto the bus drivesone line more positive and the other line more negative. This interface method providesprotection from faults, transients, and RF interference. By design, the interfaces arevirtually immune to lightning-induced transients, hot shorts, ground shorts, and RF threats.The design precludes any fault propagation (via digital audio bus) between the variousinterconnected users. At the same time, the digital audio bus interconnect structureprovides superior RF emissions characteristics, ensuring that the digital audio bus will notinterfere with sensitive receivers on-board the aircraft. The users are connected to thedata buses via a splicing arrangement (using solder rings) which experience has shown tobe extremely reliable and damage resistant. The type of cable that is specified for usemeets regulatory guidelines for flammability and smoke, and is resistant to hydraulic fluidsand fuel.

Each remote LRU contains a cluster module which, in turn, contains five digitizer chips.These are standard "off-the-shelf" chips (called CODECs - for COder/DECoder) that areused by many telephone companies. The five digitizers are sampled in sequence, theirdigital outputs are assembled into a digital data message, and the message is transmittedon the digital audio bus.

The remote COM units provide digitized COM receive audio, and the remote NAV unitsprovide digitized VOR/LOC, ADF, and MARKER BEACON audio. The NAV units also feeddiscrete digital bits (in a status byte) to enable an audio oscillator in the audio panel whenMLS or DME Morse Code Identifier audio is present. Both remote units contain additionalunassigned digitizers for future growth, one of which is frequently used for HF receivedaudio.

The two separate digital audio buses are fed to all audio panels for flight crew selection.This allows the flight crew to conveniently select and control each individual audio source.Data flow on the digital audio bus is unidirectional with a bit transmission rate of 1.0 MHz(1.0 µs/bit). Data traffic flow on the digital audio bus does not require a bus controller.The COM unit transmits a data string of approximately 60 µs every 128 µs (see Figure1-7). The NAV unit receives the COM message, synchronizes its transmitter, and then,transmits the approximately 60 µs NAV message immediately after the COM message.Should the COM unit fail, the NAV unit will go into a free run mode so as not to lose theNAV digital audio.



In each transmitted message, the preamble consists of 8 ± 1 Manchester one bits; and thesync consists of 1-1/2 bits of HIGH followed by 1-1/2 bits of LOW, which the receiver usesfor synchronization. The remaining six bytes contain eight bits each, at 1.0 µs/Bit. Thestatus byte identifies the message as COM or NAV. The digital audio panel then decodesand processes the individual bytes as appropriate to the flight crew selections.

The digital audio bus is very similar to RSB described earlier in this section. The clockfrequency is 1 MHz instead of 2/3 MHz, and the data bit assignments are different. Referto the explanation associated with Figures 1-8 thru 1-10.

128 SEC

DATA BUS

COM MESSAGE

NAV MESSAGE

C N C N C N

PREAMBLE

SYNC STATUS COM AUX1 AUX2 AUX3 AUX48 BITS EACH WORD

PREAMBLE

SYNC STATUS VORLOC

ADF8 BITS EACH WORD

AUX1 AUX2MARKER

/ /

AD-34564@

Figure 1-7. Digital Audio Data Sequence



C. Commercial Standard Digital Bus (CSDB)

The PRIMUS® 1000 system uses CSDBs for some data handling. For example, backupVOR/LOC/GS/MKR navigation display data is sent to the Radio Management Units (RMUs)from the No. 2 Navigation Unit on CSDB.

The CSDB system is made up of transmitters and receivers connected by shielded twistedpairs. Data is transmitted by a single transmitter to either a single receiver or to a group ofup to 20 receivers connected in parallel. Each CSDB carries data in one direction only.Bidirectional transmission between two Line Replaceable Units (LRUs) must beaccomplished by using two sets of transmitters, receivers, and twisted wire pair buses.

The data format is in accordance with Collins Standard 523-0772774-00611R, CommercialStandard Digital Bus (CSDB). This data bus is frequently referred to as the "CollinsPro-Line II Serial Data" bus, or simply "PL-II."

D. ARINC 429

The PRIMUS® 1000 system uses ARINC 429 data buses for most of the data handling.For example, AHRS, MADC, and EICAS data are transmitted from/to various units onARINC 429 buses.

The 429 bus system is made up of transmitters and receivers connected byshielded-twisted wire pairs. Data is transmitted by a single transmitter to either a singlereceiver or to a group of up to 20 receivers connected in parallel. Each 429 bus carriesdata in one direction only. Bidirectional transmission between two line replaceable units(LRUs) must be accomplished by using two sets of transmitters, receivers, and twisted wirepair buses.

(1) Field Definitions

ARINC 429 transmissions consist of "words" made up of 32 bits. These words aretransmitted at either 12.5 kHz (low speed) or 100 kHz (high speed) depending on thesystem. Bit number 1 is always the first bit transmitted, and bit number 32 is alwaysthe last bit transmitted. Bits 1 thru 8 are called the octal label, which identifies thetype of information contained within the word. For example, true airspeed has anoctal label of 210. In most cases, bits 9 and 10 are the source/destination identifier(SDI), which indicates the source LRU in multibox installations, by system number (1thru 4). Bits 9 and 10 may also be used as data bits in high resolution data words.Bits 11 thru 29 compose the data field. Bit 11 is the least significant bit (LSB) andbit 29 is the most significant bit (MSB). In most cases, bits 30 and 31 form the signstatus matrix (SSM), which identifies the sign and validity of the data. Like bits 9and 10 above, bits 30 and 31 may also be used as data bits in high resolution datawords. Bit 32 is used for parity.



(2) Label - Bits 1 thru 8

In the octal label, bits 1 thru 8 are used to represent numbers 0 thru 377. The eightbits are broken into two groups of three and one group of two. Each grouprepresents a digit encoded in binary with the least significant bit (LSB) having avalue of one. The octal label is transmitted with the most significant bit (MSB) of themost significant digit first. This "reversed label" characteristic is a legacy from pastsystems in which octal coding of the label field was, apparently, of no particularsignificance. Figure 1-8 shows the data bit format for octal label 274.

8 7 6 5 4 3 2 1

42 142 11 2

1111100 0

24 7

MSB

BIT NUMBER

BINARY VALUE

LSB

CHARACTER VALUE

AD-34565@

Figure 1-8. Octal Label 274

(3) Data Field - Bits 11 thru 29

Units, ranges, resolution, refresh rate, and number of significant bits for informationtransferred are encoded in either binary coded decimal (BCD), or binary (BNR)notation. Discrete information is also sent via the ARINC 429 bus. In the data field,bits 11 thru 29 are the data bits (see Figure 1-9). For some high resolution datawords, bits 9 and 10 are also data bits. Bits 30 and 31 may also be data bits.

DATA

29 28 27 26 25 24 23 22 21 20 19 18 17 16 15 14 13 12 11

LSBMSB

AD-34566@

Figure 1-9. Data Field (Bits 11 thru 29)

If bits 11 thru 29 contain data bits in a binary (BNR) format, the most significant bitof the data field represents one half of the maximum possible of the valuetransmitted. Each successive (less significant) bit represents one half of theprevious (more significant) bit. Negative numbers are encoded as the two’scomplement of positive values, with the negative sign reflected in the sign/statusmatrix.

For example, if we wish to encode a quantity whose maximum value is 2500, bitnumber 29 would represent a value of 1250, bit number 28 would represent a valueof 625, bit number 27 would represent a value of 312.5, and so on to bit number 11which would represent a value of 0.004768371541. Adding up the individual bitvalues yields the total value of the quantity being transmitted.



If bits 11 thru 29 contain data bits in a binary coded decimal (BCD) format (seeFigure 1-10), the data is grouped into four bit-bytes, each byte denoting a decimalcolumn. The 19 data bits are broken up into four groups of four bits and one groupof three bits. Each group of four can represent a number from 0 to 9; the fifth groupcan represent a number from 0 to 7. Refer to the following examples of BCD datafields. Data bit number 11 (the eleventh bit transmitted in a word) has the binaryvalue of 1. Data bits numbered 12, 13, and 14 have the arithmetic value of 2, 4, and8 respectively. Each group of bits 15 thru 29 have similarly assigned values asshown below. Using this format, decimal numbers (or characters) between 0 and 9can be assembled using combinations of these four binary values.

DATA

29 28 27 26 25 24 23 22 21 20 19 18 17 16 15 14 13 12 11

LSBMSB

8 4 2 18 4 2 18 4 2 18 4 2 14 2 1

AD-34567-R1@

DATA DATADATADATA

Figure 1-10. BCD Bit Assignments

In the data field, only those bits which are required to transmit parameter range andresolution are used, and the remaining bits are set to 0 (zero). For example, Figure1-11 shows the data word for selected course, with an octal label of 024, and a valueof 254 degrees, which only requires three characters. The remaining two charactersare filled with zeros.

29 28 27 26 25 24 23 22 21 20 19 18 17 16 15 14 13 12 11

0 1

AD-34568@

0 1 0 10 0 0 000 0 00000 1

2 5 4 XX

Octal Label: 024 Value: 254 degreesParameter: Selected Course

Figure 1-11. Selected Course Data Word

Figure 1-12 shows a DME data word which requires five characters.

29 28 27 26 25 24 23 22 21 20 19 18 17 16 15 14 13 12 11

0 1

AD-34569@

0 1 0 10 000 0 00 1

2 5 7 68

1 1 1 1 1

Octal Label: 201 Value: 257.86 NMParameter: DME Distance

Figure 1-12. DME Distance Data Word

Figure 1-13 shows a position data word requiring six characters. As can be seen,bits 9 and 10 are used, and the format is changed slightly.



%

29 28 27 26 25 24 23 22 21 20 19 18 17 16 15 14 13 12 11

AD-34570@

0 1 0 1 000 00 1

57

1 1 1 1

Octal Label: 011 Value: E 175 59.9'Parameter: Present Pos. Long.

5 9

10 9

0 1

91

1 11 0 1

Figure 1-13. Present Position Longitude Data Word

(4) Sign Status Matrix (Bits 30 and 31)

When bits 30 and 31 are being utilized for the sign status matrix (SSM) function, thebits assignments are as shown in Table 1-4.

Table 1-4. Sign Status Matrix Bit Assignments

BITMEANING

31 30

0 0 Plus, North, East, Right, To, Above

0 1 No Computed Data

1 0 Functional Test

1 1 Minus, South, West, Left, From, Below

In those data words which are BCD encoded for longitude and latitude, bits 30 and31 are both encoded to zeros for East or North, or both to ones for West or South.In addition, bits 9 and 10 are not used for SDI, but are included in the data field togive the resolution required for position.

For angular range, 0 thru 359.xxx degrees is encoded as 0 thru ±179.xxx degrees.The sign bits (30 and 31) determine the semicircle being referenced. The positiveportion of the semicircle includes 0 thru 179.xxx degrees. The negative portionincludes 180 thru 359.xxx degrees. An all-zeros configuration represents 0 and 180degrees. All ones represents 179.xxx and 359.xxx degrees. Two’s complementnotation is used for the negative half.

(5) Parity (Bit 32)

Parity is one of the simplest of all the error checking methods used in data handling.There are two basic parity configurations, ODD and EVEN. ARINC 429transmissions are always odd parity, and bit 32 is the parity bit. ARINC 429receivers are programmed to always expect an odd number of binary 1s in each32-bit word. Bit 32 is set to 1 (one) when there are an even number of binary 1s inthe word, and set to a 0 (zero) when there are an odd number of binary 1s in theword. This creates a word which always contains an overall odd number of 1s.



(6) Waveform Parameters

To be compatible with the transformer-coupled data bus, all ARINC 429 messagesare Manchester II encoded before being applied to the bus. Unlike NRZ(Nonreturn-to-Zero) data, which requires a bandwidth of dc to fc (clock frequency),Manchester encoded data is limited to the frequency range of fc/2 to fc. Also, sinceManchester data must transition in the middle of each bit period, the data clock iscontained within the data and is easily extracted at each receiver for data decoding.This feature avoids having to send a synchronous clock on separate lines along withthe data. Manchester II encoding is as follows:

ARINC 429 transmissions return to the zero voltage condition at the end of each bitperiod. As shown in Figure 1-14, a high on Line A, and a low on Line B is a binaryone. In addition, a low on Line A, and a high on Line B is a binary zero. When bothLine A and Line B are at zero volts, there is no data bit being transmitted. ARINC429 transmitters must provide a minimum dead time of four bits between messagesbecause the receivers synchronize to the transmitted data by recognizing the four-bitdead time as the synchronizing command.

/ /

/ /

1 2 3 4 5 6 7 29 30 31 32

1 1 1 1 10 0 0 0 0 0

LINE A

LINE B

AD-34572@

+50

-5

+50

-5

FOUR-BITDEAD TIME

/ /

BIT NUMBER

Figure 1-14. ARINC 429 Transmission Waveforms

Trilevel bipolar modulation consisting of HI (binary one), LO (binary zero) and NULL(no data) states are used in the transmission of data. The differential output signalvoltage across the specified output terminals (balanced to ground at the transmitter)should be as given in Table 1-5 when the transmitter is open circuit.



Table 1-5. Transmission Waveform Voltages

HI (1) NULL (V) LO (0)

Line A to Line B +10 ± 1.0 0 ± 0.5 –10 ± 1.0

Line A to Ground +5 ± 0.5 0 ± 0.25 –5 ± 0.5

Line B to Ground –5 ± 0.5 0 ± 0.25 +5 ± 0.5

The differential voltage presented at the receiver is dependent upon line length andthe number of receivers connected to a transmitter. The nominal voltage range atthe terminals is likely to be between 6.5 and 13 volts peak-to-peak. Receiver inputcommon mode voltages (Line A to Ground and Line B to Ground) are not specifiedbecause of the difficulties of defining ground with any satisfactory degree ofprecision.

The transmitter output impedance is 75 ohms balanced to ground. The receiverinput impedance is typically 8000 ohms. No more than 20 receivers (400 ohmsminimum for 20-receiver loads) should be connected to one digital data bus, andeach receiver contains isolation provisions to ensure that the occurrence of anyreasonably probable failure does not cause loss of data to the others. Bus faulttolerances for shorts and steady state voltages are designed into the transmittersand receivers.

E. RS-422

RS-422 refers to an electrical specification defined by the Electronic Industries Association(EIA). The term RS-422 is used throughout this manual to describe any data bus thatconsists of shielded twisted pairs that have not been previously described in this manual.

Examples are:

• The bus that carries data between the DA-800 Data Acquisition Units (DAUs)

• The bus that carries data from the integrated avionics computers to the display units

• The bus that carries data from the radar receiver transmitter to the display units.

F. RS-232

Like the RS-422, RS-232 also refers to an electrical specification as defined by EIA. It isused throughout this manual to describe any of the buses that are used to connect to apersonal or laptop computer. This data bus typically carries ASCII data between thecomputer and one or more of the LRUs in the PRIMUS® 1000 system.

Examples are:

• The link between the personal or laptop computer and the IAC test function

• The link between the personal or laptop computer and the DAU test function

• The link between the personal or laptop computer and the AHRS test function.



G. Serial Control Interface (SCI)

The mode, range, tilt, gain, and controller switch data is sent from the WC-650 RadarController to the WU-650 RTA. This data is sent over the serial control interface (SCI) busto the IC-600 IACs.

H. Weather Radar Picture Data (WXPD)

The weather radar picture data (WXPD) bus is a 1 MHz dedicated digital bus whichinterfaces with the RTA, MFDs and PFDs. Picture data video information is supplied fromthe RTA to the MFDs and PFDs.

I. Integrated Computer Bus (ICB)

The integrated computer bus (ICB) is a 1 MHz bus used for communication between theIC-600s. This bus operates on the HDLC hardware interface. The IC-600 uses this bus forthe following:

• Reception of data from the cross-side (xside) IC-600

• Transmission of data to the xside IC-600.

These are run-time messages. There is also a "boot" message which is controlled by theBoot software and is used for downloading new software.

This data bus provides means to communicate between the left and right IC and to loadsoftware in the IC-600.

J. SG/DU Bus

The IC-600 contains a 1 MHz picture bus that is used to transmit display formats to thePFD, MFD, and EICAS DU-870s. This bus operates on the HDLC interface.

• Each format transmission is encoded with an identifier specifying which display (PFD,MFD, and EICAS) is required to display the format

• Each IC-600 transmits data to the DU-870s at a transmission rate of every 50 ms (a 20Hz update rate), although not all the data for a complete format is sent each time.Some data is updated at slower rates, multiplexed in the 20 Hz transmissions.



(Blank Page)



SECTION 2SYSTEM DESCRIPTION

This section provides a detailed description of the general operation and cockpit displays for thePRIMUS® 1000 Integrated Avionics System. The operation and display descriptions of thePRIMUS® 1000 system are presented in subsections. Each subsection includes interface diagrams,component location diagrams, as well as outline illustrations and tables of leading particulars foreach line replaceable unit (LRU). The subsections are as follows:

• Section 2.1 - ELECTRONIC DISPLAY SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . Page 2-1-1

• Section 2.2 - ATTITUDE HEADING REFERENCE SYSTEM . . . . . . . . . . . . . . . . . . Page 2-2-1

• Section 2.3 - AIR DATA SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Page 2-3-1

• Section 2.4 - RADIO ALTIMETER SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Page 2-4-1

• Section 2.5 - WEATHER RADAR SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Page 2-5-1

• Section 2.6 - RADIO SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Page 2-6-1

• Section 2.7 - TRAFFIC/COLLISION AVOIDANCE SYSTEM . . . . . . . . . . . . . . . . . . Page 2-7-1

• Section 2.8 - FLIGHT MANAGEMENT SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . Page 2-8-1

• Section 2.9 - GLOBAL POSITIONING SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . Page 2-9-1

• Section 2.10 - FLIGHT DIRECTOR SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . Page 2-10-1

• Section 2.11 - AUTOPILOT/YAW DAMPER SYSTEM . . . . . . . . . . . . . . . . . . . . . Page 2-11-1



(Blank Page)


TABLE OF CONTENTS - SECTION 2.1



2.1 ELECTRONIC DISPLAY SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-1

2. Component Descriptions and Locations . . . . . . . . . . . . . . . . . . . . . 2-1-4

A. IC-600 Integrated Avionics Computer . . . . . . . . . . . . . . . . . . . . . . . 2-1-4

B. DU-870 Display Units . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-7

(1) Video and Dimming System . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-8

(2) System Monitor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-8

C. BL-870 Bezel Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-9

(1) IN/HPA Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-10

(2) STD (Standard) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-10

(3) BARO Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-10

D. BL-871 Bezel Controller (-831 and -851) . . . . . . . . . . . . . . . . . . . . . 2-1-11

E. DC-550 Display Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-13

(1) FULL/WX Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-15

(2) GSPD/TTG Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-15

(3) ET Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-15

(4) NAV Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-15

(5) FMS Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-15

(6) Bearing (BRG) Source Select Knobs . . . . . . . . . . . . . . . . . . 2-1-16

(7) Radio Altitude (RA) Set Knob . . . . . . . . . . . . . . . . . . . . . . . 2-1-17

(8) System TEST Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-17

F. GC-550 Guidance Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-18


(2) Course 1 (CRS 1) Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-20

(3) Heading (HDG) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-21

(4) Heading (HDG) Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-21

(5) Navigation (NAV) Button . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-21

(6) Approach (APR) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-21

(7) Bank (BNK) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-22

(8) Autopilot (AP) Engage Button . . . . . . . . . . . . . . . . . . . . . . 2-1-22

(9) Couple (CPL) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-22

(10) Yaw Damper (YD) Button . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-22

(11) Speed (SPD) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-23

(12) Vertical Speed (VS) Button . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-23

(13) Speed (SPD) Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-23

(14) Flight Level Control (FLC) Button . . . . . . . . . . . . . . . . . . . . 2-1-24

Page TC2-1-122-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.

TABLE OF CONTENTS - SECTION 2.1 (Cont)



(15) Altitude (ALT) Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-24

(16) Altitude Select (ASEL) Knob . . . . . . . . . . . . . . . . . . . . . . . . 2-1-24


(18) Course 2 (CRS 2) Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-25

G. Reversionary Panels (Embraer) . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-26

H. DA-800 Data Acquisition Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-27

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-29

A. Electronic Display System Bus Interface . . . . . . . . . . . . . . . . . . . . 2-1-29

B. PFD/MFD Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-30

(1) Bezel Controller Interface . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-30

(2) Guidance Control Unit Interface . . . . . . . . . . . . . . . . . . . . . 2-1-31

(3) Master Warning/Caution Light Interface . . . . . . . . . . . . . . . 2-1-31

C. PFD Attitude Director Indicator Operation . . . . . . . . . . . . . . . . . . . 2-1-43

(1) ADI Sphere . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-43

(2) Attitude Source Annunciations . . . . . . . . . . . . . . . . . . . . . . 2-1-47

(3) Autopilot Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-48

(4) Flight Director Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-49

(5) Vertical Deviation Display . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-55

(6) Marker Beacons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-57

(7) Radio Altitude Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-58

(8) Radio Altitude Minimums Display . . . . . . . . . . . . . . . . . . . . 2-1-58

(9) Excessive Attitude Declutter . . . . . . . . . . . . . . . . . . . . . . . 2-1-59

D. PFD Horizontal Situation Indicator Operation . . . . . . . . . . . . . . . . 2-1-60

(1) Heading Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-60

(2) Lateral Deviation Display . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-65

(3) To/From Pointer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-68

(4) Course Select/Desired Track Display . . . . . . . . . . . . . . . . . 2-1-69

(5) Drift Bug . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-70

(6) Bearing Pointers and Source Identifiers . . . . . . . . . . . . . . . 2-1-70

(7) Distance Readout . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-72

(8) Distance Identifier (FMS) . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-73

(9) Time-To-Go Readout . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-73

(10) Ground Speed Readout . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-73

(11) Elapsed Time Readout . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-73

(12) Wind Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-74

(13) Weather Radar Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-74





E. PFD Air Data Display Operation . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-79

(1) Altitude Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-79

(2) Airspeed Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-82

(3) Vertical Speed Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-86

F. PFD TCAS Display Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-89

(1) TCAS Vertical Speed Indications (VSI) . . . . . . . . . . . . . . . . 2-1-89

(2) TCAS Mode Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-92

G. PFD Miscellaneous Annunciators . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-93

(1) Air Data Computer Source Annunciator . . . . . . . . . . . . . . . 2-1-93

(2) Air Data Computer Test Annunciator . . . . . . . . . . . . . . . . . 2-1-93

(3) AHRS Test Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-93

(4) Symbol Generator Annunciator . . . . . . . . . . . . . . . . . . . . . 2-1-96

(5) Navigation Source Annunciator . . . . . . . . . . . . . . . . . . . . . 2-1-96

(6) FMS Cross-Track Mode Annunciator . . . . . . . . . . . . . . . . . 2-1-96

(7) FMS Accuracy Annunciator . . . . . . . . . . . . . . . . . . . . . . . . 2-1-97

(8) FMS Status Annunciators . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-98

(9) FMS Message Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-98

(10) Vertical Track Alert Annunciator . . . . . . . . . . . . . . . . . . . . 2-1-98

(11) ILS Approach Category Annunciator . . . . . . . . . . . . . . . . . 2-1-98

(12) Windshear Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-99

H. PFD Comparison Monitor Annunciators . . . . . . . . . . . . . . . . . . . 2-1-100

(1) Indicated Airspeed (IAS) Comparison Monitor Annunciator 2-1-100

(2) Pitch (PIT) and Roll (ROL) Attitude Comparison Monitor

Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-100

(3) Radio Altitude (RA) Comparison Monitor Annunciator . . . 2-1-100

(4) Glideslope (GS) and Localizer (LOC) Comparison Monitor

Annunciators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-100

(5) CAS Comparison Monitor Annunciator . . . . . . . . . . . . . . . 2-1-102

(6) Heading (HDG) Comparison Monitor Annunciator . . . . . . . 2-1-102

(7) Altitude (ALT) Comparison Monitor Annunciator . . . . . . . 2-1-102

I. PFD Test Mode Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-102

(1) Weight-On-Wheels Test Mode Display Formats . . . . . . . . . 2-1-102

(2) Not Weight-On-Wheels Test Mode Display Formats . . . . . 2-1-104

J. Multifunction Display Operation . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-105

(1) General Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-106

(2) MFD Bezel Menus . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-108

(3) Joystick . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-122

(4) Common MFD Map/Plan Format Data . . . . . . . . . . . . . . . . 2-1-123

(5) FMS Map/Plan Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-130





(6) MFD Map Format Display . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-133

(7) MFD Plan Format Display . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-138

(8) MFD TCAS Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-140

(9) MFD Checklist Display . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-144

(10) System Page Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-154

(11) MFD Test Mode Display . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-175

K. Engine Instrument Crew Alerting System (EICAS) Display Interface 2-1-177

L. EICAS Display Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-183

(1) Engine Instrument Section . . . . . . . . . . . . . . . . . . . . . . . . 2-1-187

(2) Cabin and APU Section . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-198

(3) Crew Alerting System Message Section . . . . . . . . . . . . . . 2-1-201

(4) Flight Control Information Section . . . . . . . . . . . . . . . . . . 2-1-209

M. EICAS Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-213

N. Reversionary Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-215

(1) Symbol Generator Reversion . . . . . . . . . . . . . . . . . . . . . . 2-1-215

(2) Sensor Reversion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-215

(3) Display Unit Reversion . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-217




FIGURE/TITLE PAGE


Figure 2-1-1. IC-600 Integrated Avionics Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-4

Figure 2-1-2. DU-870 Display Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-7

Figure 2-1-3. BL-870 Bezel Controller (-921) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-9

Figure 2-1-4. BL-871 Bezel Controllers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-11

Figure 2-1-5. DC-550 Display Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-13

Figure 2-1-6. GC-550 Flight Guidance Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-18

Figure 2-1-7. Reversionary Panels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-26

Figure 2-1-8. DA-800 Data Acquisition Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-27

Figure 2-1-9. Electronic Display System Interface Diagram . . . . . . . . . . . . . . . . . . . . . . 2-1-33

Figure 2-1-10. Pilot’s PFD/MFD Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-34

Figure 2-1-11. Copilot’s PFD/MFD Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-35





Figure 2-1-14. Pitch Attitude Scale . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-46

Figure 2-1-15. Pitch Limit Indicator Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-53

Figure 2-1-16. Primary Flight Display - HSI Display Formats (Full Heading

Compass) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-61

Figure 2-1-17. Primary Flight Display - HSI Display Formats (Partial

Heading Compass) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-62

Figure 2-1-18. Partial Heading Compass Display With Weather Radar

Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-76

Figure 2-1-19. PFD Altitude Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-80

Figure 2-1-20. PFD Airspeed Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-83

Figure 2-1-21. PFD Vertical Speed Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-87

Figure 2-1-22. PFD TCAS Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-90



Figure 2-1-24. Comparison Monitor Annunciators . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-101

Figure 2-1-25. PFD Familiarization Test Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-103

Figure 2-1-26. BL-871 MFD Bezel Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-105

Figure 2-1-27. MFD Bezel Menu Tree . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-106

Figure 2-1-28. MFD INOP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-107

Figure 2-1-29. MFD Main Menu Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-108

Figure 2-1-30. System Page Menu Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-111


TABLE OF CONTENTS - SECITON 2.1 (Cont)


FIGURE/TITLE PAGE


Figure 2-1-31. MFD Menu Display with FMS Installed . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-114

Figure 2-1-32. MFD Menu Display without FMS Installed . . . . . . . . . . . . . . . . . . . . . . . 2-1-114

Figure 2-1-33. Joystick Menu Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-116

Figure 2-1-34. Vspeeds Menu Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-118

Figure 2-1-35. Checklist Menu Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-120

Figure 2-1-36. Common MFD Map/Plan Format Data . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-124

Figure 2-1-37. Map Symbols . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-131

Figure 2-1-38. MFD Map Format Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-134

Figure 2-1-39. MFD Plan Format Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-139

Figure 2-1-40. MFD TCAS Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-141

Figure 2-1-41. MFD Checklist Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-145

Figure 2-1-42. Disclaimer Page . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-149

Figure 2-1-43. Normal Procedures Index Page . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-150

Figure 2-1-44. Waypoint Listing Page . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-151

Figure 2-1-45. Normal Checklist Page . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-152

Figure 2-1-46. Emergency Procedures Index Page . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-153

Figure 2-1-47. Electrical System Page Format - Normal Conditions . . . . . . . . . . . . . . . 2-1-155

Figure 2-1-48. Electrical System Page - Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-158

Figure 2-1-49. Hydraulic System Page Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-159

Figure 2-1-50. Hydraulic System Page - Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-161

Figure 2-1-51. Takeoff System Page Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-162

Figure 2-1-52. Takeoff System Page - Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-166

Figure 2-1-53. Environmental Control System Page Format . . . . . . . . . . . . . . . . . . . . . 2-1-167

Figure 2-1-54. Environmental Control System Page - Test Mode . . . . . . . . . . . . . . . . . 2-1-170

Figure 2-1-55. Fuel System Page Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-171

Figure 2-1-56. Fuel System Page - Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-174

Figure 2-1-57. MFD Test Mode Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-176






Figure 2-1-60. ITT Arc Default . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-190

Figure 2-1-61. ITT Arc During Engine Start . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-191

Figure 2-1-62. ITT Arc With Engine Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-191

Figure 2-1-63. EICAS Familiarization Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-214

Figure 2-1-64. Symbol Generator Reversion Mode Interface Diagram . . . . . . . . . . . . . 2-1-219


TABLE OF CONTENTS - SECITON 2.1 (Cont)


FIGURE/TITLE PAGE


Figure 2-1-65. MADC Reversion Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-220

Figure 2-1-66. AHRS Reversion Mode Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . 2-1-221

Figure 2-1-67. DAU Reversion Mode Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . 2-1-222

List of Tables

TABLE/TITLE PAGE

Table 2-1-1. IC-600 Integrated Avionics Computer Leading Particulars . . . . . . . . . . . . . . 2-1-5

Table 2-1-2. DU-870 Display Unit Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-7

Table 2-1-3. BL-870 Bezel Controller Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . 2-1-9

Table 2-1-4. BL-871 Bezel Controller Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . 2-1-12

Table 2-1-5. DC-550 Display Controller Leading Particulars . . . . . . . . . . . . . . . . . . . . . . 2-1-13

Table 2-1-6. GC-550 Guidance Control Unit Leading Particulars . . . . . . . . . . . . . . . . . . 2-1-18

Table 2-1-7. DA-800 Data Acquisition Unit Leading Particulars . . . . . . . . . . . . . . . . . . . 2-1-27

Table 2-1-10. Attitude Source Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-47

Table 2-1-11. Autopilot Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-48

Table 2-1-12. Yaw Damper Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-49

Table 2-1-13. Lateral Flight Director Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-51

Table 2-1-14. Vertical Flight Director Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-51

Table 2-1-15. Priority Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-52

Table 2-1-16. Vertical Deviation Pointer Display Colors . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-55

Table 2-1-17. Glideslope Deviation Scale . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-56

Table 2-1-18. FMS Vertical Deviation Scale (GPS Valid) . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-56

Table 2-1-19. FMS Vertical Deviation Scale (GPS Invalid) . . . . . . . . . . . . . . . . . . . . . . . 2-1-57

Table 2-1-20. Heading Source Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-64

Table 2-1-21. VOR Lateral Deviation Scale . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-65

Table 2-1-22. Localizer Deviation Scale . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-66

Table 2-1-23. FMS Lateral Deviation Scale (GPS Valid) . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-67

Table 2-1-24. FMS Lateral Deviation Scale (GPS Invalid) . . . . . . . . . . . . . . . . . . . . . . . . 2-1-67

Table 2-1-25. VOR To/From Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-68

Table 2-1-26. FMS To/From Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-68



Table 2-1-29. Weather Radar Ranges . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-75

Table 2-1-30. Color Codes for Weather Radar Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-75

Table 2-1-31. PFD WX Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-77




TABLE/TITLE PAGE


Table 2-1-32. WX Warning Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-78

Table 2-1-33. Barometric Correction Range and Resolution . . . . . . . . . . . . . . . . . . . . . . 2-1-81

Table 2-1-34. Resolution Advisory Matrix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-91

Table 2-1-35. PFD TCAS Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-92

Table 2-1-36. ADC Source Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-93

Table 2-1-37. Navigation Source Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-96

Table 2-1-38. FMS Status Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-98

Table 2-1-39. Windshear Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-99

Table 2-1-40. Menu Key Operations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-107

Table 2-1-41. MFD WX Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-127

Table 2-1-42. Heading Source Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-136

Table 2-1-43. WX and GMAP Mode Return Colors . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-137

Table 2-1-44. PFD TCAS Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-140

Table 2-1-45. Checklist Color Assignments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-146

Table 2-1-46. Waypoint Listing Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-150

Table 2-1-47. Takeoff Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-163

Table 2-1-48. Engine Takeoff Data Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-164

Table 2-1-49. N1 Indicator Dial Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-188

Table 2-1-50. ITT Engine Start Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-190

Table 2-1-51. Engine Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-197

Table 2-1-52. Ignition Annunciation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-198

Table 2-1-53. Cabin Differential Pressure Readout . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-199

Table 2-1-54. APU Turbine Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-200

Table 2-1-55. APU Exhaust Temperature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-201

Table 2-1-56. Message Inhibit Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-208

Table 2-1-57. Landing Gear Positions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-209

Table 2-1-58. Flap Positions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-210

Table 2-1-59. Pitch Trim Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-211

Table 2-1-60. DAU Reversion States . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-216

Table 2-1-61. DU Reversion States . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-1-217



SECTION 2.1ELECTRONIC DISPLAY SYSTEM

1. General

The PRIMUS® 1000 Integrated Avionics System includes an electronic display system (EDS)comprised of the following LRUs:

• Dual IC-600 Integrated Avionics Computers (IAC)

• Five DU-870 Display Units (DU)

• Two BL-870 Bezel Controllers for the PFD function

• Three BL-871 Bezel Controllers for the MFD/EICAS functions

• Dual DC-550 Display Controllers

• One GC-550 Guidance Control Unit

• Three Reversionary Control Panels (Embraer supplied)

• Two DA-800 Data Acquisition Units (DAU).

Two DUs are used to display primary flight data and are called primary flight displays (PFD).Two DUs are used as multifunction displays and are called the multifunction display (MFD). Theother DU is used to present engine indications and crew alerting messages and is called theEICAS display.

The EDS is a totally integrated system which combines the processing of primary flight displaydata with flight guidance data. This level of integration provides a number of cost and weightbenefits over traditional avionic systems and greatly simplifies the interface requirements for theflight director. The manner of integration also implies that if the EDS is operational, the flightdirector is also operational, and conversely if the EDS is failed, the flight director is also failed.This approach features all the performance advantages of display integration, flexibility,redundancy, and reliability.

Page 2-1-122-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.


The EDS displays the following information in the prime viewing area on both the pilot’s andcopilot’s PFD:

• Pitch and roll attitude

• Indicated airspeed and Mach

• Barometric altitude

• Selected alert altitude

• Heading

• Course/desired track orientation

• Vertical speed

• Flight director commands

• Mode and source annunciations.

The MFD provides the pilot or copilot with a variety of displays that are menu driven andcontrolled by six bezel-mounted pushbuttons and one rotary knob. The menu selections for thepushbuttons are shown at the bottom of the MFD. The menu selections change as a function ofwhich mode is selected for display. The MFD display formats include:

• Map display for FMS navigation

• Plan display for FMS navigation

• Weather radar display

• TCAS data

• System pages

• Electronic checklist

• Weather radar data window

• SAT/TAS/TAT data window

• Wind display.



The EICAS display is a single format display that provides full time engine and aircraft data, aswell as caution and advisory status messages for the flight crew. Engine and aircraft datadisplayed includes:

• N1 engine turbine RPMs

• N2 engine power turbine RPM

• Analog ITT scales with digital readout

• Abnormal ITT range indication

• Digital fuel flow readout

• Digital and analog readouts of main fuel quantities

• Digital and analog readouts of central fuel tank quantities

• Digital and analog readouts of oil temperature

• Digital and analog readouts of oil pressure.

• Engine low and high pressure vibration

• APU and cabin data

• Flight control information (control surfaces, landing gear, etc.)

The crew alerting system message field can display 16 messages simultaneously. Themessages fall into three categories and are differentiated by color. Warning messages are redin color (top message stack). Caution messages are amber in color (middle message stack).Advisory messages are white in color (bottom message stack). The EICAS bezel controllercontains a knob for scrolling messages into and out of the message field.



2. Component Descriptions and Locations

A. IC-600 Integrated Avionics Computer

Two IC-600 Integrated Avionics Computers are located below the cockpit floor. Figure2-1-1 shows a graphical view of the IC-600 Integrated Avionics Computer. Table 2-1-1provides items and specifications that are particular to the computer.

AD-33449@

Figure 2-1-1. IC-600 Integrated Avionics Computer



Table 2-1-1. IC-600 Integrated Avionics Computer Leading Particulars

Item Specification

Dimensions (maximum):

• Height . . . . . . . . . . . . . . . 7.62 in. (193.55 mm)

• Width . . . . . . . . . . . . . . . 4.13 in. (104.90 mm)

• Length . . . . . . . . . . . . . . 16.45 in. (418.83 mm)

Weight (maximum):

• With Autopilot . . . . . . . . . 15.5 lb (7.05 kg)

• Without Autopilot . . . . . . . 15.0 lb (6.82 kg)

Power Requirements (with autopilot):

• Continuous . . . . . . . . . . . 28 V dc, 50 W (max)

• In-Rush . . . . . . . . . . . . . . 28 V dc (0.5 sec) 200 W (max)

• Servo Power . . . . . . . . . . 28 V dc, 210 W (max)/112 W (nom)

Power Requirements (without autopilot):



Mating Connectors:

• J1, J2 . . . . . . . . . . . . . . . ITT Cannon Part No. DPX2MA-A106P-A106P-33B-0001

NOTE: Sunbank backshell (4) requiredPart No. J1560-12-2

Mounting . . . . . . . . . . . . . . . . Tray, HPN 7017095-903



The IC-600 Integrated Avionics Computer (IAC) is the primary LRU of the EDS. The pilot’sIC-600 IAC is a symbol generator, flight director, and autopilot/yaw damper computerintegrated into a single unit. The copilot’s IC-600 IAC is a flight director and symbolgenerator only. Integrating the autopilot control and flight director functions with thesymbol generator eliminates the external interfaces between these computers. All aircraftsensors and navigation sources are connected directly to the IC-600 IAC since all flightcontrol functions now reside inside this computer.

The IC-600 IAC is the focal point of information flow in the EDS. Its primary task is toconvert a variety of sensor data into digital data (word) formats for storage in memory untilthe data can be transmitted over a 1 MHz serial (EDS) bus to the PFD, MFD, and EICASdisplays. Control signals from the display and bezel controllers are used by the symbolgenerators contained within each display unit to select display format and informationsource. The system architecture also allows comparison monitoring to be performedcontinuously in the IC-600 IAC, eliminating the need for a separate comparison monitor.

Information processed in the symbol generator includes attitude (pitch and roll), heading,glideslope, localizer, course deviation, bearing (ADF, FMS and NAV), and selected air dataquantities.

The IC-600 IAC features a distributed processor architecture which utilizes independenthardware elements to perform the aircraft control and monitor functions. The architectureis designed around functional circuit card assemblies (CCAs). These separate assembliesare the power supply, analog interface, digital interface, primary CPU, and autopilot. Theautopilot CCA is not installed in the copilot’s IC-600 IAC. The two IC-600 IACscommunicate with each other via the IC bus which is a bi-directional high speed data bus.



B. DU-870 Display Units

Five DU-870 Display Units are mounted in the aircraft instrument panel. Figure 2-1-2shows a graphical view of the DU-870 Display Unit. Table 2-1-2 provides items andspecifications that are particular to the display unit.

Table 2-1-2. DU-870 Display Unit Leading Particulars

AD-29627-R1@

Figure 2-1-2. DU-870 Display Unit

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . 9.0 in. (228.60 mm)

• Width . . . . . . . . . . . . . . . . . . . . . 6.7 in. (170.18 mm)

• Length . . . . . . . . . . . . . . . . . . . . 13.53 in. (343.66 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . 25.6 lb (11.61 kg)

Power Requirements:

• Nominal . . . . . . . . . . . . . . . . . . . 28 V dc, 138 W

• Maximum . . . . . . . . . . . . . . . . . . 28 V dc, 177 W

User Replaceable Parts . . . . . . . . . . . . None

Mating Connector . . . . . . . . . . . . . . . . ITT Cannon, Part No. DPXBMA-A106-33P-0415

Mounting . . . . . . . . . . . . . . . . . . . . . . Tray, HPN 7018724-902



The DU-870 Display Unit (DU) is a large format (8-inch by 7-inch), 16-color, high resolutioncathode ray tube (CRT) and symbol generator integrated into a single LRU. The DUpresents dynamic displays to the pilot as part of the EDS. The DU symbol generator usestwo modes to generate the displays: stroke and raster. The stroke mode provides thesymbology and characters, and the raster mode provides background shades (i.e., theblue/brown sphere) and weather radar information. Stroke writing activities are directed bya vector generator which is capable of both translating and rotating characters and symbolsfor maximum display flexibility.

NOTE: The DU has a blank plate mounted to the lower bezel assembly. The PFDBL-870 or MFD/EICAS BL-871 Bezel Controller should be used as analternative.

The DUs are identical and interchangeable, except when a bezel controller is mounted tothe front of the unit. The BL-870 Bezel Controller with inclinometer is mounted to the frontof the DU when used as a PFD. The BL-871 Bezel Controller is mounted to the DU whenused as an MFD or EICAS. Provisions have been made to light the inclinometer from astandard aircraft 5-volt lighting bus. The DU wiring sends bezel controller signals to theDU rear connector.

A hold-down tray assembly holds the DU in the aircraft instrument panel. The blank plateor bezel controller must be removed to lock or unlock the DU hold-down assembly. Do notblock the center cutout in the bottom of the tray. The physical design of the DU requiresforced-air circulation for cooling its internal subassemblies. Two fans mounted on the rearof the DU provide the forced-air cooling. The fans pull air into the DU through the traycutout and ventilation holes in the bottom of the DU, where the air is then directed over thesubassemblies.

The DU also has non-volatile maintenance memory which records in-flight faults. Themaintenance memory can be read when the DU is in a factory test environment.

(1) Video and Dimming System

The DU can operate in either the raster scan or stroke writing mode. Theauto-dimming system sends a signal to the video system to control the overalldisplay intensity. In the auto-dimming system, two strategically placed ambient lightsensors generate a control signal to modulate the pilot-selected display intensity(from the dimming control on the DC-550 Display Controller).

(2) System Monitor

The DU incorporates a system monitor to provide CRT phosphor protection and asystem invalid signal to the IC-600 IAC whenever the following conditions aredetected:

• Loss of deflection in both axes

• Abnormal power supply outputs

• Improper CRT filament current.



C. BL-870 Bezel Controller

The BL-870 Bezel Controller (-921) mounts on the front of each DU-870 Display Unit that isused as a PFD. Figure 2-1-3 shows a graphical view of the BL-870 Bezel Controller.Table 2-1-3 provides items and specifications that are particular to the controller.

IN/HPA STD

BARO

AD-50627@

Figure 2-1-3. BL-870 Bezel Controller (-921)

Table 2-1-3. BL-870 Bezel Controller Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.42 in. (36.27 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.71 in. (170.51 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.10 in. (27.94 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.3 lb (0.135 kg)

User Replaceable Parts:

• Inclinometer . . . . . . . . . . . . . . . . . . . . . . . . . . . . HPN 7003115-905

• Knobs (Baro) . . . . . . . . . . . . . . . . . . . . . . . . . . . HPN 7000895-3

• Setscrew (Hex Socket, 6-32 x 3/16-inch, cup point) HPN 0455-224



The BL-870 Bezel Controller controls the following PFD display functions:

(1) IN/HPA Button

Pushing the IN/HPA button toggles the baro set digital readout on the PFD betweeninches of mercury (inHg) and HectoPascals (hPa).

(2) STD (Standard) Button

Pushing the STD button selects the standard barometric correction value for the baroset digital readout on the PFD. Barometric correction is displayed in 29.92 inHg ifinches of mercury is selected, or 1013 hPa if HectoPascals is selected.

(3) BARO Knob

The knob controls the barometric correction digital readout on the on-side PFD. Thesignal from the knob bypasses the DC-550 Display Controller and is sent directly tothe on-side MADC, which in turn provides a signal to the IC-600 IAC for displayprocessing. Rotating the knob selects a barometric correction readout in 0.01 inHgor 1 hPa increments.



D. BL-871 Bezel Controller (-831 and -851)

The BL-871 Bezel Controller (-831) mounts on the front of each DU-870 Display Unit that isused as an MFD. The BL-871 Bezel Controller (-851) mounts on the front of the DU-870Display Unit that is used as an EICAS display. Figure 2-1-4 shows a graphical view ofeach BL-871 Bezel Controller (-831 and -851). Table 2-1-4 provides items andspecifications that are particular to the controllers.

AD-50628@

BL-871 BEZEL CONTROLLER-831

BL-871 BEZEL CONTROLLER-851

Figure 2-1-4. BL-871 Bezel Controllers



Table 2-1-4. BL-871 Bezel Controller Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.42 in. (36.27 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.71 in. (170.51 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.10 in. (27.94 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.3 lb (0.135 kg)


• Knob, set . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HPN 7000895-3

• Setscrew (Hex Socket, 6-32 x 3/16-inch, cup point) HPN 0455-224

The BL-871 Bezel Controller (-831) has six bezel pushbuttons (keys) and one rotary setknob which allow the pilot or copilot to select display menu options. The menu selectionsare shown above the corresponding pushbuttons on the bottom of the MFD display format.These menu selections change as a function of which mode is selected on the MFD. Thisflexibility allows the bezel pushbuttons to control a variety of functions, while maintaining aminimum of operational complexity.

Pushing the submenu selection key causes the MFD to display that submenu. Pushing theRTN key causes the MFD to return to the top level menu, referred to as the main menu.The pilot or copilot use the rotary knob to select various map ranges when map or plandisplays are shown. Complete descriptions of the MFD menu selections are found in theoperations section of this section.

The BL-871 Bezel Controller (-851) is mounted on the front of the DU in the EICASposition. The bezel controller has a rotary knob that allows the pilot to scroll through theCAS messages on the EICAS display.



E. DC-550 Display Controller

Two DC-550 Display Controllers are mounted in the glareshield; one for the pilot and onefor the copilot. Figure 2-1-5 shows a graphical view of the DC-550 Display Controller.Table 2-1-5 provides items and specifications that are particular to the controller.

BRG BRG

NAV FMSET

OFF

NAV 1

ADF

GSPDTTG

FULLWX

FMS OFF

NAV 2

ADF

FMS

RA TEST

AD-50629-R1@

Figure 2-1-5. DC-550 Display Controller

Table 2-1-5. DC-550 Display Controller Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.25 in. (57.15 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.75 in. (146.05 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.87 in. (174.50 mm)

Weight (maximum) . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.0 lb (0.91 kg)

Power Requirements:

• Primary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 V dc, 5.0 W (max)

• Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5 V ac, 5.0 W (max)




Item Specification


• Knobs

- BRG (Setscrew A) . . . . . . . . . . . . . . . . . . . HPN 7009437

- BRG ◊ (Setscrew A) . . . . . . . . . . . . . . . . . . . HPN 7009437

- RA (Setscrew B) . . . . . . . . . . . . . . . . . . . . . . HPN 7018748-1

- Test Button HUB (Setscrew B) . . . . . . . . . . . . HPN 7009644-3

• Setscrews

- A (Multi-Spline, 2-56 x 1/8-inch, cup point) . . HPN 2500148-64

- B (Multi-Spline, 4-40 x 3/16-inch, cup point) . HPN 2500148-130

Mating Connectors:

• J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS27473E20-B35SB

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Standard Dzus Rail

The DC-550 Display Controller provides the pilot or copilot with a convenient method ofselecting the following EDS display functions:

• Selecting a bearing pointer

• Selecting a compass format for weather radar display

• Selecting groundspeed or time-to-go display

• Resetting the elapsed timer display

• Selecting the navigation source.

The display controller also provides a data acquisition function for the following remotelymounted controllers:

• PFD Bezel Controllers

• MFD Bezel Controllers

• GC-550 Guidance Controller

• Pilot and Copilot Master Caution/Warning Controller (Embraer controller)

• Reversionary Panel Controllers (Embraer controller)

• EICAS Reversion Controller (Embraer controller).

Upon receiving signals from a remote controller, the display controller transmits theacquired information to the IC-600 IAC on a two-wire digital interface bus (DC/IC). One bitis assigned on the digital bus for each pushbutton and switch input. The IC-600 IAC isconfigured through software to assign the appropriate function to each bit.



A listing of the display controller functions follows. Each function may have more than onetoggling sequence:

(1) FULL/WX Button

The pilot or copilot uses the FULL/WX button to change the PFD from a full headingcompass format to a partial heading compass format. In the full heading compassmode, 360 degrees of heading are displayed. In the partial heading compass mode,90 degrees of heading along with weather radar data. The power-up default for thisselection is FULL.

(2) GSPD/TTG Button

The pilot or copilot uses the GSPD/TTG button to display groundspeed (GPSD) ortime-to-go (TTG) in the lower right corner of the PFD. The PFD alternates betweendisplaying GSPD or TTG each time the button is pushed. If ET is currently beingdisplayed, pushing the GSPD/TTG button selects whichever parameter waspreviously displayed. The power-up default for this selection is GSPD.

(3) ET Button

The ET button allows the pilot or copilot to control an elapsed time (ET) display onthe PFD and MFD. Initial switch actuation starts the timer sequence at the previousposition. Subsequent switch actuation follows this toggle sequence:

• Reset

• Elapsed time

• Stop

• Repeat.

(4) NAV Button

The pilot or copilot uses the NAV button to select short range navigation (NAV)sources for display on the PFD. The power-up default for this selection is on-sideNAV source. The toggling sequence is as follows:

• First push: On-side NAV

• Second push: Cross-side NAV

• Repeat.

(5) FMS Button

The pilot or copilot uses the FMS button to select long range navigation (FMS)sources for display on the PFD. The power-up default for this selection is on-sideFMS source. The toggling sequence is as follows:

• First push: On-side FMS

• Second push: No effect.



If a display controller is invalid when power is applied to the system after acold-start, the following selections are automatically displayed on the PFD:

Function Left Display Right Display

Compass DisplayFormat

Full Full

FD Commands Displayed Displayed

FD Modes Inhibited Inhibited

Selected Source NAV 1 NAV 2

GSPD/TTG GSPD GSPD

(6) Bearing (BRG) Source Select Knobs

The HSI portion of the PFD can display two independent bearing pointers (BRGor BRG ). Bearing source BRG is dedicated to the sources on the left side ofthe cockpit, and BRG is dedicated to sources on the right side. The followingbearing sources can be selected for each pointer:

BRG BRG

OFFNAV1ADF1

OFFNAV2ADF2

If the display controller is invalid, the on-side NAV bearing is displayed by default.



(7) Radio Altitude (RA) Set Knob

The radio altitude control consists of a knob with a pushbutton in the center. Thepilot or copilot turns the knob to adjust the RA minimums setting which is shown asan RA digital readout on the PFD.

(8) System TEST Button

The RA set knob has a momentary action TEST button. The pilot or copilot pushesand holds the TEST button for 5 to 6 seconds while the aircraft is on the ground(weight-on-wheels) to initiate the test mode, and to do a check of the radio altimeter.

The pilot or copilot pushes the TEST button while the aircraft is in the air(weight-off-wheels) to do a check of the radio altimeter only.

NOTES: 1. The radio altimeter test is functional only if the radioaltimeter is connected to the IC-600 IAC test output. Ifconnected, the radio altimeter test can be initiated at anytime except during glideslope capture or glideslope track.

2. If the aircraft is on the ground and the TEST button is heldfor more than 5 to 6 seconds, the system enters theinitiated test mode. Refer to SECTION 7 for informationabout initiated tests.

The following test displays are shown on the PFD and MFD as long as the TESTbutton is pushed with weight-on-wheels:

• The course select, heading select, distance and GSPD/TTG digital displays arereplaced by amber dashes

• The ATT and HDG displays are flagged

• All pointers/scales are flagged

• All heading related bugs/pointers are removed

• The flight director command bars are biased from view

• The radio altimeter digital readout displays the radio altimeter self-test value

• The comparator monitor annunciates ATT, HDG, and ILS (if ILS sources areselected on both sides)

• The word TEST (in magenta) is annunciated in the lateral capture location on thetop left of the PFD

• The flight director mode annunciations are removed

• CAS MSG, RA (comparison monitor).



F. GC-550 Guidance Control Unit

The GC-550 Guidance Control Unit is mounted in the center of the glareshield. Figure2-1-6 shows a graphical view of the GC-550 Guidance Control Unit. Table 2-1-6 providesitems and specifications that are particular to the controller.

CRS 1

PUSH SYNC PUSH SYNC

HDG SPD

PUSH IAS/M

ASEL CRS 2

PUSH SYNC

FD1 HDG

BNK

APR

NAV AP

CPL

YD

SPD FD2ALT

VS

FLC

AD-50630@

Figure 2-1-6. GC-550 Flight Guidance Controller

Table 2-1-6. GC-550 Guidance Control Unit Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.25 in. (57.15 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.65 in. (295.91 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.51 in. (114.51 mm)

Weight (maximum) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.23 lb (1.01 kg)

Power Requirements:

• Panel Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5 V dc, 21.2 W (max)



Table 2-1-6. GC-550 Guidance Control Unit Leading Particulars

Item Specification


• Knobs

- CRS (Setscrew A) . . . . . . . . . . . . . . . . . . . . . . HPN 7018485-4

- HDG (Setscrew A) . . . . . . . . . . . . . . . . . . . . . HPN 7009644-1

- SPD (Setscrew B) . . . . . . . . . . . . . . . . . . . . . . HPN 7020161

- ASEL (Setscrew B) . . . . . . . . . . . . . . . . . . . . . HPN 7019971-1

- CRS 2 (Setscrew A) . . . . . . . . . . . . . . . . . . . . HPN 7018485-4

- CRS 1 PUSH SYNC (Setscrew B) . . . . . . . . . . HPN 7015342-13

- HDG PUSH SYNC (Setscrew B) . . . . . . . . . . . . HPN 7015342-12

- SPD PUSH SYNC (Setscrew B) . . . . . . . . . . . . HPN 7015342-12


• Setscrews

- A (Bristol, 4-40 x 1/8-inch, cone point) . . . . . . . HPN 2500148-128

- B (Bristol, 2-56 x 3/32-inch, cup point) . . . . . . . HPN 2500148-63

• Lamps

- Blue-White (all pushbuttons) . . . . . . . . . . . . . . HPN 7011974-2

- Clear (all pushbuttons except CPL) . . . . . . . . . HPN 7011974-6

Mating Connector:

• J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS27473E20B-35S

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Standard Dzus Rail



The GC-550 Guidance Control Unit is a 13 button controller with mode activation lightsinstalled inside each mode button. The activation light illuminates if the correspondingmode is in the arm or capture mode. The buttons provide the means of engaging theautopilot and yaw damper, selecting flight director couple, and selecting the flight directormodes for each flight director. Five rotary knobs control reference selections for verticalspeed, indicated airspeed (IAS) and Mach targets, altitude preselect, and course andheading.

The following paragraphs describe the functions of each knob and pushbutton.

(1) Flight Director 1 (FD1) Button

The primary function of the FD1 button is to turn on or off the flight director (FD)command bar display on the pilot’s PFD. There are exceptions based on the engagestatus of the autopilot as follows:

(a) Autopilot Disengaged

Pushing the FD1 button does not bring the FD command bars into view if noflight director modes are selected. Any subsequent flight director modeselection in the pitch or roll axis causes the FD command bars to be displayedon the pilot’s PFD. If the flight director mode is not selected in the pitch or rollaxis, basic attitude hold is active.

When the FD command bars are in view on both PFDs, pushing the FD1button removes the command bars from the pilot’s PFD only. When the FDcommand bars are in view on the pilot’s PFD only, pushing the FD1 buttondisengages all selected flight director modes.

If the autopilot disengages due to the A/P DISC switch, trim switch, orsecondary trim switch, the flight director remains active in the mode prior tothe autopilot disconnect. If the autopilot disengages due to an autopilot fail(AP FAIL) condition or automatic disconnects from the autopilot monitor, theFD bars on both PFDs are biased out of view and all flight director modes arecanceled regardless of the flight director engage status.

(b) Autopilot Engaged

When the autopilot is engaged, the coupled side FD command bars are alwaysdisplayed. The uncoupled side FD command bars continue to be toggled onor off with the FD1 button.

(2) Course 1 (CRS 1) Knob

The CRS 1 knob is a rotary knob with a course symbol on the knob face. This knobcontrols the course select readout on the pilot’s PFD. The knob provides theselected course to the flight director and autopilot for the VOR mode. Clockwiseknob rotation changes the selected course in one-degree increments.Counter-clockwise knob rotation changes the selected course in one-degreedecrements.



The CRS 1 knob is also an integral pushbutton which is used to synchronize course.Pushing the CRS 1 knob synchronizes the course readout on the pilot’s PFD to theaircraft’s direct-to course when VOR is the selected navigation source. The knobsignal is sent to the DC-550 Display Controller, which in turn provides a grey codesignal on the DC/IC interface bus to the IC-600 IAC for processing.

(3) Heading (HDG) Button

The primary function of the HDG button is to couple the heading select mode to boththe left-hand (LH) and right-hand (RH) flight directors. When the function isselected, the green bar on the right side of the button illuminates. This green barextinguishes when the heading select mode is deselected, or the monitor trips themode off line. The power-up default for this selection is NOT selected.

(4) Heading (HDG) Knob

The HDG knob is a rotary knob with a heading bug symbol on the knob face. Thisknob controls the heading select digital readout and the heading select bug on bothPFDs. The knob also provides the selected heading to the flight director/autopilotfor the turn direction. Clockwise knob rotation changes the selected heading inone-degree increments. Counter-clockwise knob rotation changes the selectedheading in one-degree decrements.

The HDG knob is also an integral pushbutton which is used to synchronize heading.Pushing the knob synchronizes the heading select digital readout and heading selectbug to the current aircraft heading. The knob signal is sent to the DC-550 DisplayController, which in turn provides a grey code signal on the DC/IC interface bus tothe IC-600 IAC for processing.

(5) Navigation (NAV) Button

The primary function of the NAV button is to couple the navigation mode to both theLH and RH flight directors. When the function is selected, the green bar on the rightside of the button illuminates. This green bar extinguishes when the navigationmode is deselected, or the monitor trips the mode off line. The power-up default forthis selection is NOT selected.

(6) Approach (APR) Button

The primary function of the APR button is to couple the localizer and glideslope(when tuned to an ILS frequency) or VOR approach (when tuned to VOR) to both theLH and RH flight directors, depending upon the navigation source selection throughthe DC-550 Display Controller. When the function is selected, the green bar on theright side of the button illuminates. This green bar extinguishes when the approachmode is deselected, or the monitor trips the mode off line. The power-up default forthis selection is NOT selected.



(7) Bank (BNK) Button

The primary function of the BNK button is to select a reduced maximum bank anglefor both the LH and RH flight directors when engaged in the heading select modeonly. When the function is selected, the green bar on the right side of the buttonilluminates. This green bar extinguishes when the heading select low bank mode isdeselected, or the monitor trips the mode off line. The power-up default for thisselection is NOT selected.

(8) Autopilot (AP) Engage Button

The primary function of the AP button is to engage and disengage the autopilot. Theautopilot only engages if both attitude heading reference systems are valid. Whenthe AP button is activated, the green bar on the right side of the button illuminates.Activation of the AP button also engages the yaw damper, and the green light on theright side of the YD button illuminates. Deactivation of the AP button does notdisengage the yaw damper. The green bar on the right side of the AP buttonextinguishes when the AP command is manually deselected or automaticallydisengaged. Activation of only the AP button engages the autopilot to the flightdirector that is selected by the CPL button.

(9) Couple (CPL) Button

The primary function of the CPL button is to transfer the autopilot to the coupledflight director when the autopilot is engaged. When the autopilot is not engaged, theCPL button annunciates which flight director the autopilot will couple to when itengages. The left side green triangle illuminates to indicate the autopilot will coupleto the LH flight director. The right side green triangle illuminates to indicate theautopilot will couple to the RH flight director. The power-up default for this selectionis the left side triangle.

When the autopilot is engaged, the left side or right side triangle illuminates toindicate whether the autopilot is coupled to the LH or RH flight director. Selection ofthe CPL button cancels the flight director modes independent of autopilotengagement.

(10) Yaw Damper (YD) Button

The primary function of the YD button is to engage and disengage the yaw damper.When the YD button is activated, the green bar on the right side of the buttonilluminates. Activation of the YD button also engages the autopilot, and the greenlight on the right side of the AP button illuminates. Deactivation of the YD buttondoes not disengage the autopilot. The green bar on the right side of the YD buttonextinguishes when the YD command is manually deselected or automaticallydisengaged.



(11) Speed (SPD) Button

The primary function of the SPD button is to engage the speed select mode in boththe LH and RH flight directors. When the function is selected, the green bar on theright side of the button illuminates. This green bar extinguishes when the speedselect mode is deselected, or the monitor trips the mode off line. The power-updefault for this selection is NOT selected.

(12) Vertical Speed (VS) Button

The primary function of the VS button is to engage the vertical speed select mode inboth the LH and RH flight directors. When the function is selected, the green bar onthe right side of the button illuminates. This green bar extinguishes when the verticalspeed select mode is deselected, or the monitor trips the mode off line. The verticalspeed digital reference display and bug are removed from the PFD when the verticalspeed select mode is not selected. The power-up default for this selection is NOTselected.

(13) Speed (SPD) Knob

The SPD knob is a rotary with a speed bug symbol on the face of the knob. Thisknob controls the reference digital readouts and reference bugs on both PFDs. Theknob is also an integral pushbutton which is used to toggle between IAS and Machtargets. The knob signal is sent to the DC-550 Display Controller, which in turnprovides a grey code signal on the DC/IC interface bus to the IC-600 IAC forprocessing.

(a) SPD Mode Engaged

With the speed select mode engaged, the SPD knob controls the airspeedreference digital readout on the PFDs. The knob also provides the desiredspeed command to the flight director and autopilot. Clockwise rotation of theknob changes the IAS reference in 1 knot increments and Mach in 0.01 Machincrements. Counter-clockwise rotation changes the IAS reference in 1 knotdecrements and Mach in 0.01 Mach decrements.

(b) VS Mode Engaged

With the vertical speed select mode engaged, the SPD knob controls thevertical speed reference digital readout on the PFDs. The knob also providesthe desired vertical speed command to the flight director and autopilot.Clockwise rotation of the knob changes the vertical speed reference in 100ft/min increments. Counter-clockwise rotation of the knob changes the verticalspeed reference in 100 ft/min decrements.



(c) PUSH IAS/M Button

Pushing the PUSH IAS/M button sets up the following toggle sequence whenthe speed select mode is engaged:

• If altitude is greater than 32,000 feet:

– First push: IAS

– Second push: Mach repeat

• If altitude is less than 32,000 feet:

– First push: Mach

– Second push: IAS repeat

(14) Flight Level Control (FLC) Button

The primary function of the FLC button is to send a command to both the LH and RHflight directors to provide information to change the flight level of the aircraft. Whenthe function is selected, the green bar on the right side of the button illuminates.This green bar extinguishes when the FLC mode is deselected, or the monitor tripsthe mode off line. The power-up default for this selection is NOT selected. Thetoggling sequence of the button is as follows:

• Power-up default is off

• First push: Climb/descent profile

• Second push: Off.

(15) Altitude (ALT) Button

The primary function of the ALT button is to engage the altitude hold mode in boththe LH and RH flight directors. When the function is selected, the green bar on theright side of the button illuminates. This green bar extinguishes when the altitudemode is deselected, or the monitor trips the mode off line. The power-up default forthis selection is NOT selected.

(16) Altitude Select (ASEL) Knob

The ASEL knob is a rotary knob with an altitude preselect bug symbol on the knobface. The knob controls the altitude select digital readout on both PFDs. The knobalso provides the altitude preselect to the flight director and autopilot. Clockwiserotation of the knob provides altitude selections in 100 foot increments.Counter-clockwise rotation of the knob provides altitude selections in 100 footdecrements. The knob signal is sent to the DC-550 Display Controller, which in turnprovides a grey code signal on the DC/IC interface bus to the IC-600 IAC forprocessing.



(17) Flight Director 2 (FD2) Button

The primary function of the FD1 button is to turn the flight director (FD) commandbar display on or off on the copilot’s PFD. There are exceptions based on theengage status of the autopilot as follows:

(a) Autopilot Disengaged

Pushing the FD2 button does not bring the FD command bars into view if noflight director modes are selected, Any subsequent flight director modeselection in the pitch or roll axis causes the FD command bars to be displayedon the copilot’s PFD. If the flight director mode is not selected in the pitch orroll axis, basic attitude hold is active.

When the FD command bars are in view on both PFDs, pushing the FD2button removes the command bars from the copilot’s PFD only. When the FDcommand bars are in view on the copilot’s PFD only, pushing the FD2 buttondisengages all selected flight director modes.

If the autopilot disengages due to the wheel master switch, cooley hat switch,or secondary trim switch, the flight director remains active in the mode prior tothe autopilot disconnect. If the autopilot disengages due to an autopilot fail(AP FAIL) condition or automatic disconnects from the autopilot monitor, theFD bars on both PFDs are biased out of view and all flight director modes arecanceled regardless of the flight director engage status.

(b) Autopilot Engaged

When the autopilot is engaged, the coupled side FD command bars are alwaysdisplayed. The uncoupled side FD command bars continue to be toggled onor off with the FD2 button.

(18) Course 2 (CRS 2) Knob

The CRS 2 knob is a rotary knob with a course symbol on the knob face. This knobcontrols the course select readout on the copilot’s PFD. The knob also provides theselected course to the flight director and autopilot for the VOR mode. Clockwiseknob rotation changes the selected course in one-degree increments.Counter-clockwise knob rotation changes the selected course in one-degreedecrements.

The CRS 2 knob is also an integral pushbutton which is used to synchronize course.Pushing the CRS 2 knob synchronizes the course select digital readout on the pilot’sPFD to the aircraft’s direct-to course when VOR is the selected NAV source. Theknob signal is sent to the DC-550 Display Controller, which in turn provides a greycode signal on the DC/IC interface bus to the IC-600 IAC for processing.



G. Reversionary Panels (Embraer)

Three Embraer supplied reversionary panels are installed in the cockpit. Two reversionarypanels are mounted in the instrument panel, and the EICAS reversionary panel is mountedin the pedestal. The reversionary panels control the selection of the IAC, ADC, AHRS, andDAU that is being used (system reversion). The reversionary panels also allow the pilot tocontrol which DU is displaying a particular format (DU reversion). Figure 2-1-7 shows agraphical view of the reversionary panels.

These reversionary panels provide reversion control in three areas:

• Symbol generator reversion

• Sensor reversion

• Display unit reversion.

AD-50631-R1@

REVERSIONARY PANEL

NORMEICASPFD

AHRS SG

EICAS REV

DAU 1 DAU 2

MFD ADC

Figure 2-1-7. Reversionary Panels



H. DA-800 Data Acquisition Unit

Figure 2-1-8 shows a graphical view of the DA-800 Data Acquisition Unit (DAU).Table 2-1-7 provides items and specifications that are particular to the unit.

AD-29414@

Figure 2-1-8. DA-800 Data Acquisition Unit

Table 2-1-7. DA-800 Data Acquisition Unit Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . 7.62 in. (193.55 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . 3.59 in. (91.19 mm)

• Length (from rear connector) . . . . . . . 12.62 in. (320.55 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . 10.0 lb (4.535 kg)

Power Requirements . . . . . . . . . . . . . . . . 28 V dc, 30 W (max)

User Replaceable Parts . . . . . . . . . . . . . . None

Mating Connector . . . . . . . . . . . . . . . . . . Tri-Star Part No. TR2P106P106P-0001(210)

Mounting . . . . . . . . . . . . . . . . . . . . . . . . Tray, HPN 7014882-901



The DA-800 DAU is the central data collection point for the engine instrument display(EICAS). The EICAS display provides information on all required engine parameters, aswell as engine related systems such as fuel. The EICAS display also provides an area forthe crew advisory system (CAS) to display warning, caution, and advisory messages. TheDAU also provides interface to external functions that need data, such as the flight datarecorder (FDR) and ground proximity warning system (GPWS).

One DAU is provided for each side of the aircraft. Left and right engine and aircraftsensors are connected to DAU No. 1. Right engine and aircraft sensors are connected toDAU No. 2. Each DAU has dual operating channels with independent power supplies foroperational redundancy.

The DAU digitizes various analog and discrete signals and sends the data to the IC-600IAC via an ARINC 429 data bus.



3. Operation

A. Electronic Display System Bus Interface

The IC-600 IAC contains the symbol generator (SG) function for the EDS (seeFigure 2-1-9). The IAC SG receives all inputs on high speed ARINC 429 data buses andthe radio system bus (these buses are not shown) and a digital bus from the DC-550Display Controller. The SG function processes data from the various sources and sendsoutputs to the five DU-870 Display Units (DU).

The display controller receives analog signals consisting of switch position, pushbutton,and knob rotation data from the GC-550 Guidance Control Unit, reversionary panels, andthe MFD and EICAS bezel controllers. The display controller digitizes these analog signalsfor transmission to the IAC over a two-wire digital interface bus (DC/IC) running at a 9600baud rate.

The IAC contains three 1 MHz bus inputs for DU status and wrap-around data. The single1 MHz bus output connects to the DUs to provide display format information. The IACcontains two display electronic interfaces (DEIs) which interface to the DU output bus. TheDEIs use data from the I/O processor to generate the proper display formats for the DUs.Each DEI has the ability to generate two independent formats, and output them on the DUoutput bus. Display controller information, along with reversionary controller switchpositions is also used by the DEI to determine which display format is placed on the DUoutput bus.

Each IAC receives ARINC 429 and EDS wraparound data from the DUs. Under normaloperation (no reversion), the pilot’s IAC monitors the wraparound bus activity counter of DUNo. 1 for DU validity. The copilot’s IAC monitors the wraparound bus activity counter of DuNo. 3 and DU No. 5 for DU validity. In addition, each DU displaying PFD informationprovides a wraparound bus for displayed pitch, roll, altitude, indicated airspeed (IAS), andbarometric correction. The DU displaying the engine information provides a wraparoundbus for N1, N2, and ITT data from each engine. The PFD and EICAS information receivedfrom the wraparound bus is compared to the appropriate sources for DU validity. If thewraparound monitor is tripped, a crew alerting system (CAS) message is annunciated

When one of the IACs fails, the remaining IAC supplies four independent display formats.The remaining DEIs provide the following formats:

• Independent PFD for each pilot

• MFD format

• EICAS format.

The 1 MHz bus that supplies the DUs with data from the IAC provides a number ofadvantages to the system.

• Minimum interconnect wiring

• Each DU can input up to four different data buses

• Each data bus can transmit two independent display formats

• Selection of data buses and formats is done through select discretes at the DU.



B. PFD/MFD Interface

The DC-550 Display Controller is the data concentrator for:

• BL-870 Bezel Controller (PFD)

• BL-871 Bezel Controller (MFD)

• GC-550 Guidance Control Unit

• Master Warning/Caution Control Panel (Embraer controller).

Control signals from the bezel controllers and guidance control unit are used to select thedisplay formats and information sources for the EDS. Upon receiving signals from thesecontrollers, the display controller transmits the acquired information to the IC-600 IAC onthe DC/IC interface bus. One bit is assigned on the bus for each pushbutton and selectknob input.

The IAC is configured through software to assign the appropriate function to each bit. TheSG function converts digital inputs from the display controller into the graphic formats thatare required by the PFD and MFD. The IAC then outputs the display formats on DU busNo. 1 or No. 2 to the PFD or MFD.

(1) Bezel Controller Interface

The bezel controllers for the PFD and MFD control many of the display functions.Each PFD bezel controller contains two pushbuttons and a BARO set knob. EachMFD bezel controller contains six pushbutton menu keys and a rotary knob for menucontrol. See Figure 2-1-10 for the pilot’s PFD/MFD interface and Figure 2-1-11 forthe copilot’s PFD/MFD interface.

Pushing the IN/HPA button on the PFD bezel controller toggles the barometriccorrection setting between inches of mercury (inHg) and HectoPascals (hPa).Rotating the BARO set knob on the bezel controller adjusts the inHg or hPa digitaldisplay on the on-side PFD. The BARO set knob produces a grey code which issent directly to the AZ-840 MADC to provide baro correction. The MADC thenprovides a signal to the IAC for display processing. Rotating the knob selects inHgin 0.01 inHg increments, or in 1 hPa increments. Pushing the STD button on thePFD bezel controller selects the standard barometric correction of 29.92 inHg or1013 hPa for the PFD display.

The MFD bezel controller pushbuttons allow the pilot or copilot to select a menufunctions being displayed on the MFD. Pushing the menu key associated with asubmenu function causes the MFD to display that submenu. Pushing the RTN keycauses the MFD to return to the top level menu (main menu). The pilot or copilotuses the rotary knob to select various map ranges when map or plan displays areshown. If weather radar is selected for display, the rotary knob has no control sincethe weather radar range is controlled by the dedicated WC-650 Radar Controller.The pilot uses the rotary knob on the EICAS bezel controller to scroll through thedisplayed CAS messages.



The WU-650 Receiver/Transmitter/Antenna (RTA) generates a serial data interfacewhich provides scan-converted data of the weather radar picture. The RTA transmitsthis data on a 1 MHz bus to the left and right PFDs and MFDs. The data beingtransferred includes color coding of the radar data, X-Y memory location, and systemtiming. Weather radar data is displayed on the PFD when the WX format is selectedby the display controller, and on the MFD when WX is selected from the menuoption.

A dimming controller, consisting of five 10k ohm potentiometer, interfaces with eachDU to independently dim each unit. The DU supplies a voltage and groundreference while receiving the wiper voltage over a twisted, shielded triple wire set.

(2) Guidance Control Unit Interface

Thirteen pushbuttons on the GC-550 Guidance Control Unit enable the pilot toengage the autopilot and yaw damper, select flight director couple, and select theflight director modes for each flight director. Mode activation lights are installedinside each mode pushbutton. When a mode is engaged in the IAC, the computerproduces a corresponding annunciator output. This annunciator output is sent to theguidance control unit to illuminate the appropriate mode light. See Figure 2-1-12 forthe guidance controller interface diagram.

The guidance control unit has five select knobs that interface with the displaycontroller. The select knobs provide grey code inputs to the display controller toselect the aircraft course, heading, airspeed, and altitude. The course and headingknobs are also integral pushbuttons which are used to synchronize course select andheading select to the current aircraft course and heading. A push to change buttonlocated inside the speed knob allows the pilot to select either an IAS target or Machtarget as the airspeed reference. The PFD displays the airspeed reference (IAS orMach) in a digital readout. The PFD uses the speed reference to control movementof the speed bug on the speed scale. The display controller converts each graycode into a digital format and transmits the data to the IAC on the DC/IC interfacebus.

(3) Master Warning/Caution Light Interface

The crew warning panel provides the logic for illuminating and flashing the twomaster warning/caution lights which are installed in the glareshield. Each masterwarning/caution light has two parts contained in a pushbutton switch. The upper partof the switch is a red warning annunciator, and the lower part is an amber cautionannunciator. An active CAS warning message causes the IAC to generate a discreteoutput for the red warning annunciator. The annunciator then illuminates andflashes. An active CAS caution message causes the IAC to generate a discreteoutput for the amber caution annunciator. The annunciator then illuminates andflashes.

Pushing the master warning/caution light produces a message acknowledge discretewhich the IAC uses in its CAS message processing logic. Pushing the light alsoproduces a reset signal for the display controller, which subsequently causes the IACto turn off the red or amber annunciators.



(Blank Page)



DU NO. 1 - PFD 1130J1

3536

3839

BUS NO. 2 IN

BUS NO.1 IN

IC-600 NO. 1 190J2B

1011

14

EDS BUS OUT

AD-49858-R1@

6679

ARINC 429WRAPAROUND OUT

DU NO. 2 - MFD 1131J1

3536

3839

BUS NO. 2 IN

BUS NO. 1 IN

3132

WRAPAROUNDBUS OUT

DU NO. 3 - EICAS132J1

3536

3839

BUS NO. 2 IN

BUS NO. 1 IN

3132

WRAPAROUNDBUS OUT

DU NO. 4 - MFD 2C131J1

3536

3839

BUS NO. 2 IN

BUS NO. 1 IN

3132

WRAPAROUNDBUS OUT

DU NO. 5 - PFD 2C130J1

3536

3839

BUS NO. 2 IN

BUS NO. 1 IN

6679

ARINC 429WRAPAROUND OUT

EDS BUS NO. 1WRAPAROUND IN

(TERMINATED)

EDS BUS NO.2WRAPAROUND IN

EDS BUS NO. 3WRAPAROUND IN(UNTERMINATED)

ARINC 429WRAPAROUND IN

190J2A

15

1617

2728

4546

IC-600 NO. 2 C190J2B

1011

14

EDS BUS OUT


(TERMINATED)


EDS BUS NO. 3WRAPAROUND IN(UNTERMINATED)

ARINC 429WRAPAROUND IN

NO.2 TERM

C190J2A

15

1617

2728

4546

47

127OHMS

127OHMS

115J1

1516

C190J2A

DC/IC BUS IN

15

16DC/IC BUS IN

DC-550 NO. 2

3435

DC/IC BUS OUT

115J1DC-550 NO. 1

3435

DC/IC BUS OUT

190J2A

47NO. 2 TERM

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

H

L

Figure 2-1-9. Electronic Display System Interface Diagram

Pages 2-1-33/3422-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.


SIGNAL GROUND

SIGNAL GROUND

WX BUS NO. 1

2841 DLS OUT

(H)(L)

2942

DLS OUT

(H)(L)ALS IN

2841

(H)(L)

SET KNOB COMMON 63 BEZEL COMMON23

62

47

80

130J1

131J1

AIRCRAFT DIMMING CONTROL

BRIGHTNESS POT

0-5 V DC DIMMING 2324

115J1

IN/HPA

SET KNOB

DU NO. 2

115J1

AZ-850 MADC NO. 1

7778

(GND / OPEN)

POWER GROUND

28 V DCINPUT POWER

SIGNAL GROUND

ALS IN

80

DU-870 NO. 1 (PFD)

BRIGHTNESS POT

24LIGHTING COMMON

23 0-5 V DC DIMMING0-5 VDC EDGE LIGHTING

PB IN/HPA

SIGNAL GROUNDPB STD BARO

63BEZEL COMMON

DC-550 DISPLAY CONTROLLER NO. 1

9J1

0 - 5 V DC EDGE LIGHTING

PUSH BUTTON NO. 1PUSH BUTTON NO. 2PUSH BUTTON NO. 3PUSH BUTTON NO. 4PUSH BUTTON NO. 5

(H)(L)

2125

4647505160

(H)(L)


SET KNOB

13141

462521

(H)(W)(L)

(L)(H)BARO

KNOB

64

WU-650/WU-870 RTA59J1

(H)PICTURE BUS

BUS TERMINATING RESISTORNO. 1

7778(L)

(H)WX BUS(L)

LEFTgh

11413

(H)(W)(L)

LIGHTING COMMON

(H)(L)

2942

PUSH BUTTON NO. 6

(H)(L)

66

747571

AD-50635@

(H)(L)

BAROCORRECTION

BARO STD SYNC

PUSH BUTTON NO. 6

2122

7847484956

DU-870 NO. 2 (MFD)

POWER GROUND104105106

28 V DCAM DCBUS 1A

28 V DCINPUT POWER

103102101 28 V DC

AM DCBUS 2A

104105106

101102103

Figure 2-1-10. Pilot’s PFD/MFD Interface Diagram



SIGNAL GROUND

SIGNAL GROUND

WX BUS NO. 1

2841 DLS OUT

(H)(L)

2942

DLS OUT

(H)(L)ALS IN

2841

(H)(L)

SET KNOB COMMON 63 BEZEL COMMON23

62

47

80

C130J1

C131J1

AIRCRAFT DIMMING CONTROL

BRIGHTNESS POT

0-5 V DC DIMMING 2324

C115J1

IN/HPA

SET KNOB

DU NO. 2

C115J1

AZ-850 MADC NO. 2

7778

(GND / OPEN)

POWER GROUND

28 V DCINPUT POWER

SIGNAL GROUND

ALS IN

80

DU-870 NO. 5 (PFD)

BRIGHTNESS POT

24LIGHTING COMMON

23 0-5 V DC DIMMING0-5 VDC EDGE LIGHTING

PB IN/HPA

SIGNAL GROUNDPB STD BARO

63BEZEL COMMON

DC-550 DISPLAY CONTROLLER NO. 2

C9J1

0 - 5 V DC EDGE LIGHTING


(H)(L)

2125

4647505160

(H)(L)


SET KNOB

13141

462521

(H)(W)(L)

(L)(H)BARO

KNOB

64

WU-650/WU-870 RTA59J1

(H)PICTURE BUS

BUS TERMINATING RESISTORNO. 1

7778(L)

(H)WX BUS(L)

RIGHTks

11413

(H)(W)(L)

LIGHTING COMMON

(H)(L)

2942

PUSH BUTTON NO. 6

(H)(L)

66

747571

AD-50636@

(H)(L)

BAROCORRECTION

BARO STD SYNC

PUSH BUTTON NO. 6

2122

7847484956

DU-870 NO. 4 (MFD)

POWER GROUND104105106

28 V DCAM DCBUS 1A

28 V DCINPUT POWER

103102101 28 V DC

AM DCBUS 2A

104105106

101102103

Figure 2-1-11. Copilot’s PFD/MFD Interface Diagram



IC-600 IAC NO. 1

FD 1 ON PB 48

HDG PB IN 22

NAV PB IN 24APR PB IN 25ALT PB IN 26

SPD PB IN 27FLC PB IN 28

VS PB IN 30LAMP DRVR OUT 4 FLC 37

LAMP DRVR OUT 3 FD 1 ON 36MASTER CAUTION LAMP OUT 35

190J1B

COUPLE ANNUNC OUT 99HDG MODE ANNUNC OUT 106

BNK ANNUNC OUT 98NAV MODE ANNUNC OUT 105APR MODE ANNUNC OUT 104

ALT MODE ANNUNC OUT 103IAS MODE ANNUNC OUT 102

VS MODE ANNUNC OUT 101

190J2A

AP ENGAGE SEL PB IN 85YD ENGAGE SEL PB IN 83

COUPLE PB IN 55BANK PB IN 56

AP ENGAGE ANNUNC OUT 86

YD ENGAGE ANNUNC OUT 84MASTER WARNING LAMP OUT 20

190J2B


789101112

1314151617

PUSHBUTTON NO. 1 (FD 1)PUSHBUTTON NO. 2 (AP ENG)PUSHBUTTON NO. 3 (YD ENG)PUSHBUTTON NO. 4 (CPL)PUSHBUTTON NO. 5 (HDG)PUSHBUTTON NO. 6 (BNK)

PUSHBUTTON NO. 7 (NAV)PUSHBUTTON NO. 8 (APR)PUSHBUTTON NO. 9 (ALT)PUSHBUTTON NO. 10 (SPD)PUSHBUTTON NO. 11 (FLC)

18 PUSHBUTTON NO. 12 (RESERVED)19 PUSHBUTTON NO. 13 (VS)

2324252627

2829303132

LAMP NO. 2 (AP ENG)LAMP NO. 3 (YD ENG)LAMP NO. 4 (CPL L)LAMP NO. 5 (HDG)

LAMP NO. 6 (BNK)

LAMP NO. 7 (NAV)LAMP NO. 8 (APR)LAMP NO. 9 (ALT)LAMP NO. 10 (SPD)LAMP NO. 11 (FLC)

33 LAMP NO. 12 (FD 1)34 LAMP NO. 13 (VS)

1 PILOT MODE GROUND

4 LIGHTING COMMON3 PANEL LIGHTING 5 VOLTS (H)

5 ANNUNCIATOR LIGHTINGDAY/NIGHT

SWITCH

SIGNALGROUND

0-5 V DCEDGELIGHTING

PUSHBUTTON NO. 14 (FD 2)LAMP NO. 1 (FD 2)

LAMP NO. 11 (FLC)

PUSHBUTTON NO. 4 (CPL)PUSHBUTTON NO. 5 (HDG)PUSHBUTTON NO. 6 (BNK)

PUSHBUTTON NO. 7 (NAV)PUSHBUTTON NO. 8 (APR)PUSHBUTTON NO. 9 (ALT)

PUSHBUTTON NO. 10 (SPD)PUSHBUTTON NO. 11 (FLC)

PUSHBUTTON NO. 13 (VS)

LAMP NO. 5 (HDG)

LAMP NO. 6 (BNK)

LAMP NO. 7 (NAV)LAMP NO. 8 (APR)LAMP NO. 9 (ALT)

LAMP NO. 10 (SPD)

LAMP NO. 15 (CPL R)LAMP NO. 13 (VS)

202232

101112

131415161719

2627

28293031

3634

HDG PB IN22

NAV PB IN24APR PB IN25ALT PB IN26

SPD PB IN27FLC PB IN28

VS PB IN30FD 2 PB IN48LAMP DRVR FD 2 ON36LAMP DRVR FLC37MASTER CAUTION LAMP OUT35

11J1 11J1 C190J1B

C190J2A

VS MODE ANNUNC OUT101

HDG MODE ANNUNC OUT106BNK ANNUNC OUT98NAV MODE ANNUNC OUT105APR MODE ANNUNC OUT104

ALT MODE ANNUNC OUT103IAS MODE ANNUNC OUT102

COUPLE ANNUNC OUT99

555620

C190J2B

COUPLE PB INBANK PB INMASTER WARNING LAMP OUT

MASTERWARNING

MASTERCAUTION

+28 V DC

A

B

MASTERWARNING

MASTERCAUTION

+28 V DC

AD-50740, SH1@

C

D

IC-600 IAC NO. 2

Figure 2-1-12 (Sheet 1). Guidance Control Unit Interface Diagram



DC-550DISPLAY CONTROLLER NO. 1

707168

910

11

12

SET KNOB NO. 1SET KNOB NO. 2

SET KNOB COMMON

SET KNOB NO. 1SET KNOB NO. 2

SET KNOB COMMON1314

15SET KNOB NO. 1SET KNOB NO. 2

SET KNOB COMMON1617

18SET KNOB NO. 1SET KNOB NO. 2

SET KNOB COMMON1920

MASTER CAUTION RESET 60

61MASTER WARNING RESET

115J1

626365

11J1

CRS 1 PUSH TO SYNCHDG PUSH TO SYNCSPD (IAS/MACH SEL)

444546

SET KNOB NO. 1SET KNOB NO. 2SET KNOB COMMON

CRS 1

474849


HDG

505152


ALT

535455


SPD

565758


CRS 2

66 CRS 2 PUSH TO SYNC

GC-550FLIGHT GUIDANCE CONTROLLER

CRS 1

HDG

ALT

SPD

73AHRS REVERSION

74MADC REVERSION

AHRS REV 1 AHRS REV 2

MADC REV 1 MADC REV 2

TO XPDR 1 TO XPDR 2

DC-550DISPLAY CONTROLLER NO. 2

PUSHBUTTON (PUSH TO SYNC CRS 2)707168

910

11

12

PUSHBUTTON (PUSH TO SYNC HDG)



1314

15 SET KNOB NO. 1SET KNOB NO. 2SET KNOB COMMON

1617

18 SET KNOB NO. 1SET KNOB NO. 2SET KNOB COMMON

1920

C115J1

CRS 2

HDG

ALT

SPD

73 AHRS REVERSION

74 MADC REVERSION

PUSHBUTTON (SPD IAS/MACH SEL)

PUSHBUTTON (PUSH TO SYNC CRS 1)PUSHBUTTON (PUSH TO SYNC HDG)

PUSHBUTTON (SPD IAS/MACH SEL)

60 MASTER CAUTION RESET

61 MASTER WARNING RESET

AD-50740, SH2@

A

B

C

D

Figure 2-1-12 (Sheet 2). Guidance Control Unit Interface Diagram



C. PFD Attitude Director Indicator Operation

The attitude director indicator (ADI) display symbology uses a truncated sphere format fordisplaying standard attitude information. The AHRS provides the pitch and roll attitudeinputs. The ADI display formats are shown in Figure 2-1-13, sheets 1 and 2.

(1) ADI Sphere

The ADI sphere presents the aircraft body-axis pitch and roll attitude with respect tothe horizon. Within the ADI sphere is sky/ground shading, an aircraft symbol, a pitchtape, and a roll attitude pointer. A roll attitude scale is displayed on the top of theADI sphere.

(a) Sphere Shading

The ADI sphere is divided into two halves separated by a white horizon line,with raster shading in both halves. The upper raster shading is cyan torepresent the sky. The lower raster shading is brown to represent the ground.The raster moves with respect to the aircraft symbol such that the pitch scaleshows the actual pitch of the aircraft and the roll scale shows the actual roll ofthe aircraft. An eyebrow of raster shading is displayed for pitch attitudesgreater than 17.5° and less than -17.5°. The eyebrow can be used as areference for rapidly determining the location of ground. The eyebrow is notdisplayed for invalid pitch or roll attitude.

(b) Attitude Reference Aircraft Symbol

The attitude reference aircraft symbol serves as a stationary representation ofthe aircraft. Aircraft pitch and roll attitudes are displayed by the relationshipbetween the fixed aircraft and the moveable ADI sphere. The miniatureaircraft symbol is flown to align the command cue to the aircraft symbol inorder to satisfy flight director computed steering commands. Two types ofaircraft symbols are presented: single cue and cross pointer. The pilot canselect either symbol for display from the DC-550 Display Controller.

1 Single Cue Aircraft Symbol

The yellow single cue aircraft symbol is comprised of a triangle positionedwith one side parallel to the horizontal bottom of the DU. An indentation ornotch exists along the bottom edge of the symbol. Outrigger boxes resideat the left and right edge of the ADI sphere.

2 Cross Pointer Aircraft Symbol

The yellow cross pointer aircraft symbol is comprised of three parts. Twolateral bars are spaced the width of the single cue notch. A small squareis positioned on the center of the ADI sphere between the bars.



HORIZONLINE

260

235

245

001HDG

359

E

W

21

12

30

15

33

S

24

200

260

280

220

.410 M

VOR1

ADF2

YDIASHDG

200 RA

14500

14500

GS

3

1000

1

2

3000

0

1

2

3

25

LOC

29.92 IN

20

20 20

20

10

10 10

10

0020

80240

1

9

AD-51187@

GSAP

MIN

143

N3

6

CRS

TGTTTG5MIN

13.1 NM

NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS

VERTICALDEVIATIONSCALE, POINTERAND LABEL

RA MINIMUMSANNUNCIATOR

RADIO ALTITUDEDIGITALREADOUT

LOW ALTITUDEAWARENESS

RA MINIMUMDIGITALREADOUT

SINGLE CUEFLIGHTDIRECTORCOMMAND BARS

YAW DAMPERSTATUS

ANNUNCIATORPITCH LIMITER

INDICATORLOW

BANK ARC

AUTOPILOTSTATUS

ANNUNCIATOR

ROLLATTITUDEPOINTER

ROLLATTITUDE

SCALE

PITCHATTITUDE

SCALE

SINGLE CUEAIRCRAFT

SYMBOL

ATTITUDESPHERE

350

Figure 2-1-13 (Sheet 1). Primary Flight Display - ADI Display Formats




M

MIN1

260

25

260

24240

220

200

280

LOC IASHDGAP YD

10 10

20 20

10 10

20 20

145 00

14500

14320

8000

GS

1

9 R2

FMS

14000

MAX

SPD

AOA

ATT2

CROSS POINTERFLIGHT DIRECTORCOMMAND BARS

CROSS POINTERAIRCRAFT SYMBOL

MARKER BEACON

IMO

ARMEDVERTICALFD MODE

ANNUNCIATOR

CAPTUREDVERTICALFD MODE

ANNUNCIATOR

FLIGHTDIRECTOR

COUPLE ARROW

CAPTUREDLATERALFD MODE

ANNUNCIATOR

ARMEDLATERALFD MODE

ANNUNCIATOR

AUTOPILOTOVERSPEED

WARNINGANNUNCIATOR

ATTITUDE SOURCEANNUNCIATOR

ATT1ATT2

AD-51188@

001HDG

359

E

W

21

12

30

15

33

S

24

.410 M

VOR1

ADF2

200 RA

3

1000

1

2

3000

0

1

2

3

25

29.92 IN

N3

6

CRS

TGTTTG5MIN

13.1 NM

350

Figure 2-1-13 (Sheet 2). Primary Flight Display - ADI Display Formats



(c) Pitch Attitude Scale

The ADI sphere contains a pitch attitude tape that indicates the pitch attitudeof the aircraft. The tape is linear and displays pitch between zero and ±90°.The tape moves down for positive inputs and up for negative inputs. The taperotates about the center of the ADI sphere and moves with the roll pointer toshow roll information. The PFD removes the pitch tape scale, pitch tapereference mark labels, chevrons, and the horizon line for an invalid pitchattitude.

The pitch attitude tape is shown in Figure 2-1-14, and is described as follows:

• Reference marks are provided at the following positions:

– positive 0, 5, 10, 15, 20, 25, 30, 40, 60, and 90

– negative 5, 10, 15, 20, 25, 30, 45, 60, and 90

• Identifying digits are provided at the following positions:

– positive 10°, 20°, 30°, 40°, 60°, and 90°– negative 5°, 10°, 20°, 30°, 45°, 60°, and 90°

• Red excessive pitch chevrons are displayed at 45° and 65° pitch up, and35°, 50°, and 60° pitch down.

30 30

10 10

20 20

40

60

90

10 10

20 20

30

45

60

90

AD-51189@

Figure 2-1-14. Pitch Attitude Scale



(d) Roll Attitude Scale

The PFD displays the white roll attitude scale along the top edge of the ADIsphere. The roll attitude scale displays ±60° of roll although the roll pointer isnot limited to the roll attitude scale. Tick marks are provided along the top ofthe ADI sphere at the following angles:

• ±10°, ±20°, ±30°, and ±60°

• The ±30° tick mark is twice as long as the other tick marks.

Inverted triangles are located at 0° and ±45° of roll attitude.

(e) Roll Attitude Pointer

The white roll attitude pointer is a triangle symbol that moves along the insideof the ADI sphere. The pointer displays ±180° of roll. The pointer movescounterclockwise for positive inputs, and clockwise for negative inputs. ThePFD removes the pointer for an invalid roll attitude.

(2) Attitude Source Annunciations

The pilot or copilot can select the attitude source through pushbutton switches on thetwo reversionary panels. The switches select either AHRS No. 1 or AHRS No. 2 asthe source of attitude data on the corresponding PFD. Table 2-1-10 defines theannunciations and colors associated with each attitude source selection.

Table 2-1-10. Attitude Source Annunciations

SG ReversionSelection

AHRS ReversionSelection

AttitudeSensor

AnnunciationColor

Pilot Copilot Pilot Copilot Pilot Copilot

Normal Normal Normal Normal None NONE N/A

Normal Normal Reversion Normal ATT2 ATT2 Amber

Normal Normal Normal Reversion ATT1 ATT1 Amber

Normal Normal Reversion Reversion ATT2 ATT1 Amber

Reversion Normal N/A Normal ATT2 ATT2 Amber

Reversion Normal N/A Reversion ATT1 ATT1 Amber

Normal Reversion Normal N/A ATT1 ATT1 Amber

Normal Reversion Reversion N/A ATT2 ATT2 Amber



(3) Autopilot Display

The autopilot/yaw damper processor in the pilot’s IC-600 IAC provides the autopilotinformation for display on both PFDs. The autopilot/yaw damper processor alsoprovides the status messages that are displayed on the PFD.

(a) Autopilot Status Annunciator

The PFD displays autopilot annunciations above the ADI sphere, andpositioned horizontally just left of center. The autopilot annunciation isdisplayed in a four-character field which provides autopilot system statusindications for the pilot. Flashing autopilot annunciations flash at a rate of 1second on/0.5 second off. Autopilot annunciations are defined in Table2-1-11.

Table 2-1-11. Autopilot Annunciations

Annunciation Color Activating Conditions State

AP Green Autopilot Engage = Engaged Steady

AP Amber Autopilot Engage = Disengaged (See Note 1)Autopilot Fail = Not Fail

5 secondflash

AP Red Cat II Active and Autopilot Engage =Disengaged (See Note 1)

5 secondflash

AP Red Autopilot Engage = Disengaged (See Note 1)Autopilot Fail = Fail

5 secondflash; thensteady

AP TEST(See Note 2)

Amber Autopilot Test = Test Steady

TCS White Touch Control Steering (TCS) = Engaged Steady

TKNB Amber Turn Knob Out of Detent = Engaged Steady

NOTES:

1. This is a transition from engaged to not engaged.

2. The TEST annunciations are displayed in the yaw damper field for acombined display of AP TEST.



(b) Yaw Damper Status Annunciator

The PFD displays yaw damper annunciations above the ADI sphere, andpositioned horizontally just right of center. The yaw damper annunciation isdisplayed in a four-character field which provides yaw damper system statusindications for the pilot. Flashing yaw damper annunciations flash at a rate of1 second on/0.5 second off. Yaw damper annunciations are defined in Table2-1-12.

Table 2-1-12. Yaw Damper Annunciations

Annunciation Color Activating Conditions Display

YD Green Yaw Damper Engage = Engaged Steady

YD Amber Yaw Damper Engage = Disengage (See Note 1)Yaw Damper Fail = Fail

5 secondflash; thensteady

YD Amber Yaw Damper Engage = Disengaged (See Note 1)Yaw Damper Fail = Not Fail

5 secondflash

TEST(See Note 2)

Amber Autopilot Test = Test Steady

NOTES:

1. This is a transition from engaged to not engaged.

2. The TEST annunciation is displayed in the yaw damper field for a combineddisplay of AP TEST.

(4) Flight Director Display

Flight director (FD) data is processed by the flight director function in the IAC anddisplayed on both PFD formats. The flight director function provides the followingdisplay information:

• Left-right flight director couple arrow

• Lateral flight director mode (capture and arm) annunciations

• Vertical flight director mode (capture and arm) annunciations

• Flight director command bars.

(a) Flight Director Couple Arrow

The PFD displays the green flight director couple arrow at the top center of thedisplay format. The arrow points either left or right in the direction of the flightdirector to which the autopilot is coupled if engaged.



(b) Flight Director Mode Annunciators

The PFD displays flight director modes annunciations to the left and right ofthe flight couple arrow. The information displayed represents armed andcaptured lateral and vertical flight director modes.

Lateral mode annunciations are displayed on the upper left side of the PFD.Vertical mode annunciations are displayed on the upper right side of the PFD.Captured lateral and vertical modes are annunciated in green in four-characterfields. Armed lateral and vertical modes are annunciated in white infour-character fields located to either side of the respective captured mode.

Flight director modes with both an arm and capture state have a modetransition defined as follows:

• A white transition box surrounds the captured mode for eight secondsimmediately after the transition from armed to captured has occurred if themode was previously armed.

• The transition box does not surround a captured mode if the mode was notpreviously armed.

• If a new mode is captured during a previous mode transition, the eightsecond timer is reset.

• A transition box surrounds a default mode when a transition to defaultoccurs automatically.

• Transitions are as follows:

Lateral Transitions Vertical Transitions

BC arm to BC capture ASEL arm to ASEL capture

LOC arm to LOC capture ASEL capture to ALT capture

VAPP arm to VAPP capture GS arm to GS capture

VOR arm to VOR capture WSHR arm to WSHR capture

Any mode to ROL

Any mode to PIT

Table 2-1-13 lists the lateral mode annunciations for valid flight directorconditions. Table 2-1-14 lists the vertical mode annunciations for valid flightdirector conditions. Table 2-1-15 lists the priority logic if more than twovertical modes are armed simultaneously.



Table 2-1-13. Lateral Flight Director Modes

Annunciation Description Notes

BC Back Course Mode Arm or Capture

LNAV Long RangeNavigation Mode

Capture mode only

HDG Heading Select Mode Capture mode only

LOC Localizer Mode Arm or Capture

ROL Basic Roll Mode Capture mode only

VAPP VOR Approach Mode Arm or Capture

VOR VOR Navigation Mode Arm or Capture

Table 2-1-14. Vertical Flight Director Modes

Annunciation Description Notes

ALT Altitude Hold Mode Capture mode only

ASEL Altitude Preselect Mode Arm or Capture

CLB Climb Mode Capture mode only

DES Descent Mode Capture mode only

GA Go-Around Mode Capture mode only

GS Glide Slope Mode Arm or Capture

IAS Indicated Airspeed Mode Capture mode only

MACH Mach Mode Capture mode only

PIT Basic Pitch Mode Capture mode only

TO Takeoff Mode Capture mode only

WSHR Windshear Mode Arm or Capture

VS Vertical Speed Mode Capture mode only



Table 2-1-15. Priority Logic

Vertical Arm Field

GS Highest

WSHR

ASEL Lowest

(c) Flight Director Command Bars

The flight director provides pitch and roll commands to the pilot or copilot bydisplaying single cue or cross pointer command bars on the ADI sphere. Theflight director function provides the pitch command, roll command, andin-view/validity flags. When the flight director sets the in-view flags, thecommand bars are shown within the ADI sphere. The PFD removes the flightdirector command bars for the following conditions:

• The flight director is invalid

• The in view flags are not set by the flight director function

• The FD button on the GC-550 Guidance Control Unit has deselected thecommand bars

• The command bars are biased out of view for an invalid navigation setup.

1 Single Cue Command Bars

The magenta single cue command bars are composed of two elongatedtriangles. These command bars are only displayed when a lateral orvertical mode has been selected. The command bars represent computedlateral and vertical steering commands. To satisfy the steering command,the pilot flies the single cue aircraft symbol to the command bars.

For zero degrees of flight director pitch and roll command, the single cuecommand bars rest on top of the aircraft symbol. Increasing values ofpitch command cause an upward movement of the single cue commandbars. Increasing values of roll command cause a clockwise movement ofthe single cue command bars.

2 Cross Pointer Command Bars

The magenta cross pointer command bars are composed of twostroke-written cues, one vertical and the other horizontal. These commandbars are displayed when a lateral or a vertical mode has been selected.The lateral cross pointer command bar moves for lateral commands, andthe vertical cross pointer command bar moves for vertical commands. Tosatisfy the steering command, the pilot flies the cross pointer aircraftsymbol to the appropriate command bar.



In the zeroed condition, the cross pointer cues intersect at the nose of thecross pointer aircraft symbol. Increasing values of pitch command causean upward movement of the lateral cue. Increasing values of rollcommand cause a movement of the vertical cue to the right.

(d) Pitch Limit Indicator

A pitch limit indicator (PLI) symbol is drawn directly on the ADI sphere torepresent the current stickshaker angle of attack. If the aircraft symbol meetsor exceeds the PLI, the aircraft goes into stickshaker. The PFD displays thePLI symbol parallel to the aircraft symbol on the ADI sphere. The PLI symbolis located either above or below the aircraft symbol at the value of attitudedegrees indicated by the flight director PLI margin. A positive PLI marginplaces the PLI symbol above the aircraft symbol, while a negative PLI marginplaces the PLI symbol below the aircraft symbol. At a PLI margin of zerodegrees, the PLI symbol is located on the top edge of the single cue aircraftsymbol, or the center of the cross pointer aircraft symbol.

The PLI symbol is displayed in its normal color (green) if the PLI margin isgreater than five degrees. If the PLI margin is less than or equal to fivedegrees but greater than two degrees, then the PLI symbol is displayed in thecaution exceedance color (amber). The PLI symbol is displayed in thewarning exceedance color (red) if the PLI margin is less than or equal to twodegrees. The PFD removes the PLI symbol if the pitch limit indication fromthe flight director is invalid. Figure 2-1-15 shows the PLI symbol in relation toboth the single cue aircraft symbol and the cross pointer aircraft symbol on theADI sphere.

20

20 20

20

10

10 10

10 10 10

20 20

10 10

20 20

AD-51190@

Figure 2-1-15. Pitch Limit Indicator Display



(e) Low Bank Arc

The green low bank arc provides an indication on the roll scale of the flightdirector bank limit. The PFD displays the low bank arc immediately above theADI sphere, extending ±14° from the zero-degree roll tick triangle. The PFDremoves the low bank arc if the flight director is invalid.

(f) Autopilot Overspeed Warning Annunciator

The autopilot overspeed warning provides an indication that the autopilot isoperating in an overspeed condition. This warning is normally provided whenthe autopilot is coupled to vertical speed or a VNAV flight director mode.When the overspeed condition is met, the flight director transitions to a speedmode until the overspeed condition drops. The PFD displays an amber MAXSPD annunciation vertically next to the ADI sphere to indicate an overspeedwarning condition.

(g) Autopilot Underspeed Warning Annunciator

The autopilot underspeed warning provides an indication that the autopilot isoperating in an underspeed condition. This warning is normally provided whenthe autopilot is coupled to vertical speed or a VNAV flight director mode.When the underspeed condition is met, the flight director transitions to a speedmode until the underspeed condition drops. The PFD displays an amber MINSPD annunciation vertically next to the ADI sphere to indicate an underspeedwarning condition.



(5) Vertical Deviation Display

The PFD displays vertical deviation information to the right of the ADI sphere. Thewhite vertical deviation scale indicates glide slope (GS) deviation or long rangenavigation (FMS) deviation. The vertical deviation scale also indicates invalidvertical deviation.

The scale contains a horizontal line representing the zero index position which alignswith the horizon line of the ADI sphere when the aircraft is at zero pitch and roll.Two whites circles, refereed to as dots, are spaced evenly above and below thehorizontal line. A truncated, left-pointing triangle represents the deviation pointer.The pointer moves vertically along the scale to indicate the vertical deviation of theaircraft. Increasing values of short range navigation (SRN) glide slope or long rangenavigation (LRN) vertical deviation produces a downward movement of the pointer.The color of the pointer is specified in Table 2-1-16.

Table 2-1-16. Vertical Deviation Pointer Display Colors

Navigation Source Vertical Deviation Display

Pilot SelectedSource

Copilot SelectedSource

Pointer Color onPilot PFD

Pointer Coloron Copilot PFD

SRN1 SRN2 Green Green

SRN2 SRN1 Yellow Yellow

SRN1 FMS Green Magenta

SRN2 FMS Yellow Magenta

FMS SRN1 Magenta Yellow

FMS SRN2 Magenta Green

(a) Vertical Deviation (VOR)

The PFD automatically removes the vertical deviation scale, label, and pointerwhen the selected navigation source is VOR and the tuned to localizer (TTL)indicates a tuned-to-VOR frequency.

(b) Vertical Deviation (GS)

When the selected navigation source is VOR/LOC, and TTL indicates atuned-to-localizer frequency, the PFD displays the vertical deviation scale andpointer, plus a white GS label above the scale. Each dot in the scalerepresents the vertical deviation from beam center. The deviation pointermoves vertically along the scale as specified in Table 2-1-17. For invalidglideslope vertical deviation, the PFD removes the pointer and draws a red Xthrough the center of the vertical deviation scale.



Table 2-1-17. Glideslope Deviation Scale

PointerPosition

VerticalDeviation Input

DDM

2nd Dot Up +150 +0.175

1st Dot Up +75 +0.0875

Zero Index 0 0

1st Dot Down -75 -0.0875

2nd Dot Down -150 -0.175

(c) Vertical Deviation (FMS)

When the selected navigation source is FMS, the PFD displays the FMSvertical deviation scale and pointer, plus a white FMS label above the scale.Each dot represents 250 feet from vertical path center when not in theapproach phase, and 75 feet when in the approach phase. The deviationpointer is driven based on the data presented in Table 2-1-18 when the GPSmode is valid. The deviation pointer is driven based on the data presented inTable 2-1-19 if the GPS mode is invalid. For invalid FMS vertical deviation,the PFD removes the scale, label, and pointer from the display.

Table 2-1-18. FMS Vertical Deviation Scale (GPS Valid)

Bar Position GPS ApproachVertical Deviation

(feet)(GPS Mode =

Approach)

GPS Terminal AreaVertical Deviation

(feet)(GPS Mode =

Terminal Area)

EnrouteVertical Deviation

(feet)(GPS Mode =

Enroute)

2nd Dot Up -150 -500 -500

1st Dot Up -75 -250 -250

Zero Index 0 0 0

1st Dot Down +75 +250 +250

2nd Dot Down +150 +500 +500



Table 2-1-19. FMS Vertical Deviation Scale (GPS Invalid)

PointerPosition

FMS ApproachVertical Deviation (feet)

(FMS Scaling =Approach)

EnrouteVertical Deviation (feet)(FMS Scaling = Enroute)

2nd Dot Up -150 -500

1st Dot Up -75 -250

Zero Index 0 0

1st Dot Down +75 +250

2nd Dot Down +150 +500

(d) Excessive Vertical Deviation (GS)

The excessive deviation monitor is activated when a CAT2 condition is met.When the excessive deviation monitor trips, the vertical deviation indicationprovides an alert to the pilot as follows:

• The vertical deviation pointer turns amber

• The vertical deviation scale turns amber and flashes.

(6) Marker Beacons

The PFD displays the marker beacons next to the altitude tape. Any active markerbeacon flashes continuously on the display. An active marker beacon is onlydisplayed if the selected navigation is an SRN. A white box outlines the activemarker beacon annunciation. For invalid marker beacon data, the PFD removes therespective marker beacon annunciation.

The navigation radio provides the input for each marker beacon. The markerbeacons have no priority. If the bit is set for a marker beacon, then that beacon isdisplayed. Each marker beacon is displayed if all three bits are set. The markerbeacons are displayed as follows:

• Inner: annunciated by flashing a white I inside the outline box

• Middle: annunciated by flashing an amber M inside the outline box

• Outer: annunciated by flashing a cyan O inside the outline box.



(7) Radio Altitude Display

The PFD displays radio altitude in a digital readout at the bottom of the ADI sphere,and as a shading color change on the altitude tape. The radio altimeter provides theinputs for the radio altitude display.

(a) Radio Altitude Digital Readout

The radio altitude digital readout has a range from -20 feet to +2,550 feet.The readout resolution is in 5 foot increments from -20 to +200 feet, and in 10foot increments from +200 to +2,550 feet. A white box outlines the digitalreadout. The PFD removes the readout and outline box for radio altitudesgreater than 2,550 feet.

Readout digits are displayed in green. When the TEST button on the DC-550Display Controller is pushed, the readout displays a radio altitude valid valueof 100 feet. For an invalid radio altitude, the PFD replaces the readout digitswith an amber RA annunciation.

(b) Low Altitude Awareness Indicator

The PFD displays the low altitude awareness indication as a shading colorchange on the altitude tape. Color shading provides an indication of groundwith respect to the current altitude. At a radio altitude of zero feet, the entirelower half of the altitude tape is shaded brown. The shading rises linearly forradio altitudes from +550 feet down to zero feet. A yellow horizontal line isdrawn across the altitude tape at the shading transition. The PFD removesthe horizontal line for radio altitudes less than or equal to 60 feet.

Any of the following conditions cause the PFD to remove the low altitudeawareness shading and horizontal line:

• An invalid radio altitude

• An invalid barometric altitude.

(8) Radio Altitude Minimums Display

The PFD displays radio altitude (RA) minimums in a digital readout, and as anannunciation on the PFD. The radio altitude minimums is invalid if the DC-550Display Controller is invalid.

(a) RA Minimums Digital Readout

The PFD displays the digital readout below and to the right of the ADI sphere.A white RA label annunciates the readout (an RA minimums value of zeroremoves the RA label). The RA minimums values in the readout are set byturning the RA set knob on the DC-550 Display Controller. The readout rangehas a range from 5 feet to 999 feet. The readout resolution is in 5 footincrements from -20 to +200 feet, and in 10 foot increments from +200 to +990feet. Any RA minimums values above 990 feet are forced to 999 feet. Thepower-up default is 200 feet. The readout digits are displayed in cyan. Threeamber dashes (---) replace the digits when RA minimums is invalid.



(b) RA Minimums Annunciator

The RA minimums annunciator indicates that the radio altitude is within acertain range of RA minimum digital readout. A white box outlines theannunciator field when RA minimums becomes armed and radio altitude iswithin 100 feet of the RA minimums digital readout. The RA minimumsannunciator becomes captured when the radio altitude is less than or equal tothe RA minimums readout. An amber MIN annunciation is displayed inside thebox, flashes for 10 seconds, and then remains steady to indicate captured.The PFD removes the RA minimums annunciator when RA minimums or radioaltitude are invalid.

(9) Excessive Attitude Declutter

When an excessive attitude situation occurs, the PFD removes certain symbology todeclutter the display format. An excessive attitude situation is declared when eitherof the following conditions are met:

• Roll attitude is greater than +65° or less than -65°

• Pitch attitude is greater than +30° up or less than -20° down.

The PFD removes the following symbology from the display format for an excessiveattitude situation:

• All failure flags for the items listed

• Decision height digital readout and label

• Flight director couple arrow

• Flight director command bars

• Heading, radio altitude, localizer, glide slope, and ILS comparator monitorannunciations.

• Low bank limit arc

• Marker beacons

• Radio altitude digital readout, box, and raster mark

• Selected airspeed bug, digital readout and outline

• Selected altitude bug, digital readout and outline

• Vertical and lateral FD mode annunciations

• Vspeed bugs and readouts.

The PFD restores the removed symbology when the following conditions are met:

• Roll attitude is less than +63° or greater than -63°

• Pitch attitude is less than +28° up or greater than -18° down.



D. PFD Horizontal Situation Indicator Operation

The PFD presents the horizontal situation indicator (HSI) as a compass display. TheAHRS provides the magnetic heading inputs. The DC-550 Display Controller provides theselected heading inputs.

(1) Heading Display

The heading display includes a compass card with a stationary aircraft symbol, aheading digital readout for the partial heading compass display, and a heading selectdigital readout and bug. The pilot or copilot uses the FULL/WX pushbutton on theDC-550 Display Controller to select the full heading compass display or partialheading compass display. Figure 2-1-16 shows the full heading compass display.Figure 2-1-17 shows the partial heading compass display.

(a) Full Heading Compass Scale

The full heading compass display provides a 360-degree compass card. Whitelong tick marks are displayed at integral multiples of 10 degrees, and whiteshort tick marks at the intermediate five-degree positions. Digits and cardinalabbreviations are spaced around the inside of the compass card at 30°increments. Eight numeric identifiers (3, 6, 12, 15, 21, 24, 30, and 33) arelocated at 30, 60, 120, 150, 210, 240, 300, and 330 degrees, respectively.Four cardinal abbreviations (N, E, S, and W) are shown at 0, 90, 180, and 270degrees.

A white lubber line is provided at the apex of the compass card as a pointer tothe current heading value. Seven tick marks related to the lubber line arepositioned at 45-degree intervals around the outside of the compass card.

A white stationary aircraft symbol is displayed in the center of the compasscard. The aircraft symbol indicates the aircraft’s position relative to the actualheading and selected heading. The compass card rotates either clockwise orcounterclockwise around the aircraft symbol depending on the current heading.For an invalid heading, the PFD displays a red HDG FAIL annunciation abovethe stationary aircraft symbol.




LATERALDEVIATION SCALE

3000

260

235

245

001HDG

359

E

W

21

12

30

15

33

S

24

200

260

280

220

.410 M

VOR1

ADF2

YDIASHDG

MAG2

14500

14500

GS

3

1000

1

2

0

1

2

3

25

LOC

29.92 IN

20

20 20

20

10

10 10

10

0020

80240

1

9

AD-51191@

GSAP

143

N3

6

CRS

TGTTTG5MIN

13.1 NM

350

COURSESELECT/DESIREDTRACK POINTER

TO/FROMPOINTER

LATERALDEVIATION

BAR

BEARINGPOINTER NO. 2

HEADINGSELECT BUG

HEADING SELECTDIGITAL READOUT

BEARINGPOINTER NO. 1

TIME-TO-GOREADOUT

DISTANCEREADOUT

CURRENT HEADINGLUBBER LINE

COURSE/DESIREDTRACK READOUT

CRS DTK

BEARING POINTERINDENTIFIERS

ADF 1/2FMS 1/2VOR 1/2

HEADING SOURCEANNUNCIATOR

DG 1DG 2

MAG 1MAG 2TRU 1TRU 2

HEADINGCOMPASS

DRIFT BUG

Figure 2-1-16. Primary Flight Display - HSI Display Formats (Full Heading Compass)




LOC

R1

2

.410 M 29.92 IN

DTK359

HDG001

0

1

23

32

133N

3603000

1000

VOR1ADF2

12.5HDR

25

GSPD245KTS

MAG2

260

235

245

200

260

280

220

YDIASHDG 14500

GS

20

20 20

20

10

10 10

10

2401

9

GSAP

14000

14500

0020

80143

DISTANCEIDENTIFIER

(TO WAYPOINT)

AD-51192@

FMS

KDVT

HEADINGDIGITALREADOUT

WIND ANGLEARROW

WIND SPEEDREADOUT

DME HOLDANNUNCIATOR

GROUND SPEEDREADOUT

350

Figure 2-1-17. Primary Flight Display - HSI Display Formats (Partial Heading Compass)



(b) Partial Heading Compass Scale

The partial heading compass display provides a ±45 degree arc with a digitalreadout for an accurate interpretation of the heading. White long tick marksare displayed at 10 degree intervals, and white short tick marks atintermediate five degree intervals. Digits and cardinal abbreviations arespaced around the inside of the compass arc at 30-degree increments. Eightnumeric identifiers (3, 6, 12, 15, 21, 24, 30, and 33) are located at 30, 60,120, 150, 210, 240, 300, and 330 degrees. Four cardinal abbreviations (N, E,S, and W) are shown at 0, 90, 180, and 270 degrees. All digits and cardinalabbreviations rotate with the compass.

A white stationary aircraft symbol is displayed at the centerpoint of the partialcompass arc. The aircraft symbol indicates the aircraft position relative tomagnetic north. The partial compass arc rotates either clockwise orcounterclockwise around the aircraft symbol depending on the current heading.Increasing values of heading cause a counterclockwise rotation of thecompass arc. For invalid heading, the MFD displays red characters HDGFAIL, centered above the stationary aircraft symbol.

(c) Heading Digital Readout

The PFD displays a digital readout of the current heading when the partialheading compass display is selected. The readout has a range from 001° to360° with a resolution of one degree. Leading zeros are provided for headingvalues less than 100.

A white pointer box serves as a place holder for the readout digits, and as apointer to the current heading value. The readout digits are displayed ingreen. For an invalid heading, three amber dashes (---) replace the digits.

(d) Heading Select Digital Readout

The PFD displays the heading select value in units of degrees in a digitalreadout. The readout has a range from 001° to 360° with a resolution of onedegree. The pilot or copilot turns the HDG knob on the GC-550 GuidanceControl Unit to select a value for the readout. When the HDG knob is pushed,the readout slews to the coupled-side current heading display.

A white HDG label is displayed above the readout. The readout digits aredisplayed in white. For an invalid heading select, three amber dashes (---)replace the readout digits.

(e) Heading Select Bug

The PFD displays a cyan heading select bug on the outside of the compassarc. The bug is capable of 360 degrees of motion, with a resolution ofone-degree increments. Increasing values of heading select cause aclockwise movement of the bug. The pilot or copilot turns the HDG knob onthe GC-550 Guidance Control Unit to control the bug movement. When theHDG knob is pushed, the heading select bug slews to the coupled-side currentheading display.



If the heading select bug is off-scale, an out-of-view arrow is displayedpointing in the direction of the bug. For an invalid heading select, the PFDremoves the bug and arrow.

(f) Heading Source Annunciation

The PFD displays the heading source annunciations directly above theheading compass scale. Each heading source annunciation is based on theposition of the AHRS button on the reversionary panels. Table 2-1-20 lists theheading source annunciations that are available for display, along with theirdisplay colors. The reversion annunciation is displayed in a caution color ifboth the pilot and copilot are displaying the same AHRS data, or if they areboth displaying cross-side AHRS data.

Table 2-1-20. Heading Source Annunciations



Pilot PFD Display Copilot PFD Display

Norm Norm Norm Norm DG1(W)

" " TRU1(W)

DG2(W)

" " TRU2(W)

Norm Norm Norm Rev DG1(A)

MAG1(A)

TRU1(A)

DG1(A)

MAG1(A)

TRU1(A)

Norm Norm Rev Norm DG2(A)

MAG2(A)

TRU2(A)

DG2(A)

MAG2(A)

TRU2(A)

Norm Norm Rev Rev DG2(A)

MAG2(A)

TRU2(A)

DG1(A)

MAG1(A)

TRU1(A)

Rev Norm N/A Norm DG2(A)

MAG2(A)

TRU2(A)

DG2(A)

MAG2(A)

TRU2(A)

Rev Norm N/A Rev DG1(A)

MAG1(A)

TRU1(A)

DG1(A)

MAG1(A)

TRU1(A)

Norm Rev Norm N/A DG1(A)

MAG1(A)

TRU1(A)

DG1(A)

MAG1(A)

TRU1(A)

Norm Rev Rev N/A DG2(A)

MAG2(A)

TRU2(A)

DG1(A)

MAG2(A)

TRU2(A)

NOTES:

(A) indicates amber(W) indicates white" " indicates that there is no display when magnetic is displayed and both pilots have selected

their normal heading source.



(2) Lateral Deviation Display

The PFD displays lateral deviation in the center of the heading compass display.The white lateral deviation scale indicates course deviation for VOR/LOC navigation,as well as desired track deviation for FMS navigation. At the center of the deviationscale is the aircraft symbol. The aircraft symbol represents the aircraft’s actualposition relative to the selected course.

The deviation scale has two white circles (referred to as dots) spaced evenly on bothsides of the aircraft symbol, and positioned perpendicular to the courseselect/desired track pointer. The entire scale is displayed at all times for full headingcompass and partial heading compass displays. The pointer indicates a "fly to"condition, and the deviation scale rotates with the pointer.

The lateral deviation scale contains a deviation bar which represents the centerlineof the selected VOR or localizer course. Increasing values of lateral deviation ordesired track cause the bar to move right. Decreasing values of course deviation ordesired track cause the bar to move left. The PFD removes the lateral deviation barwhen heading is invalid for any selected navigation source.

(a) VOR Lateral Deviation

When the selected navigation source is VOR, the deviation bar is driven basedon the data in Table 2-1-21. The deviation bar is displayed in the same coloras the course/desired track readout. Beyond the second scale dot, thedeviation bar continues to move for inputs up to ±12°, but at a reducedsensitivity. For an invalid course deviation, the PFD removes the deviation barand draws a red X through the center of the lateral deviation scale.

Table 2-1-21. VOR Lateral Deviation Scale

Bar Position VOR Computed Deviation(Course - VOR Bearing)

2nd Dot right -10°

1st Dot right -5°

Zero Index 0°

1st Dot left +5°

2nd Dot left +10°



(b) Localizer Lateral Deviation

When the selected navigation source is an ILS, the deviation bar is drivenbased on the data in Table 2-1-22. The deviation bar is displayed in the samecolor as the course/desired track readout and pointer. Beyond the secondscale dot, the deviation bar continues to move for inputs up to ±185 µA, but ata reduced sensitivity. For invalid localizer deviation, the PFD removes thedeviation bar and draws a red X through the center of the lateral deviationscale.

Table 2-1-22. Localizer Deviation Scale

Bar Position Localizer Deviation (µA/ddm)

2nd Dot right +150/0.155

1st Dot right +75/0.0775

Zero Index 0/0

1st Dot left -75/-0.0775

2nd Dot left -150/-0.155

(c) Excessive Lateral Deviation (LOC)

The excessive deviation monitor activates when a CAT2 condition is met.When the excessive deviation monitor trips, the lateral deviation bar turnsamber and flashes to alert to the pilot or copilot.



(d) FMS Lateral Deviation

When the selected navigation source is an FMS, the PFD displays thecross-track deviation scale and bar. The deviation bar is positionedperpendicular to the deviation scale. The deviation bar is driven by the lateralscale factor from the FMS when the scale factor equation is valid. Thedeviation bar is driven based on data in Table 2-1-23 when the scale factorequation is invalid and the GPS mode from the FMS is valid. The deviationbar is driven based on data in Table 2-1-24 if both the scale factor equationand GPS mode are invalid.

The deviation bar is displayed in the same color as the course/desired trackreadout and pointer. For an invalid cross-track deviation, the PFD removesthe deviation bar and draws a red X through the center of the lateral deviationscale. The PFD only removes the deviation bar for an invalid heading.

Table 2-1-23. FMS Lateral Deviation Scale (GPS Valid)

Bar Position GPS ApproachCross-Track

Deviation (NM)(GPS Mode =

Approach)

GPS TerminalArea

Cross-TrackDeviation (NM)(GPS Mode =

Terminal Area)

EnrouteCross-Track

Deviation (NM)(GPS Mode =

Enroute)

2nd Dot Right -0.3 -1.0 -5.0

1st Dot Right -0.15 -0.5 -2.5

Zero Index 0 0 0

1st Dot Left +0.15 +0.5 +2.5

2nd Dot Left +0.3 +1.0 +5.0

Table 2-1-24. FMS Lateral Deviation Scale (GPS Invalid)

PointerPosition

FMS ApproachVertical Deviation (feet)

(FMS Scaling =Approach)

EnrouteVertical Deviation (feet)(FMS Scaling = Enroute)

2nd Dot Right -1.0 -5.0

1st Dot Right -0.5 -2.5

Zero Index 0 0

1st Dot Left +0.5 +2.5

2nd Dot Left +1.0 +5.0



(3) To/From Pointer

The PFD displays the white To/From pointer near the head or tail of the lateraldeviation bar in the HSI display. The pointer indicates whether the aircraft is flyingto or from a VOR station or waypoint, depending on the selected navigation mode.The pointer rotates with the course select/desired track pointer.

(a) VOR To/From Indications

When the selected navigation source is a VOR, the PFD displays the To/Frompointer based on the information listed in Table 2-1-25. The PFD removes thepointer for invalid to/from data.

Table 2-1-25. VOR To/From Indications

Position VOR Bearing Selected Course

To Less than or equal to 88°

From Greater than or equal to 92°

Not Displayed Between 88 and 92°

(b) FMS To/From Indications

When the selected navigation source is a FMS, the PFD displays the To/Frompointer based on information listed in Table 2-1-26. The PFD removes thepointer from the display for invalid to/from data.

Table 2-1-26. FMS To/From Indications

Position LRN Input

To To/From = To

From To/From = From

Not Displayed To/From = Neither To nor From



(4) Course Select/Desired Track Display

The PFD displays course select information when the selected navigation source isan SRN source. The pilot or copilot uses the course set knob on the GC-550Guidance Control Unit to control the course selection. The PFD displays desiredtrack information when the selected navigation source is FMS. The FMS providesthe desired track information which can not be selected by the pilot or copilot.

The PFD displays course select and desired track in a shared digital readout, and asa pointer. Increasing values of course select or desired track cause an increase inthe readout value and a clockwise movement of the pointer. The N cardinal on theheading compass is the 360-degree position for course select and desired track.

(a) Course Select Readout (VOR, ILS)

The course select digital readout has a range from 001° to 360° with aresolution of one degree. The readout digits are displayed in yellow when thenavigation source is a VOR, or in green when the navigation source is an ILS.A CRS label annunciates the digital readout when course select is set. TheCRS label is displayed in the same color as the digital readout.

The digital readout slews to the current navigation source bearing when thecourse knob on the GC-550 Guidance Control Unit is pushed and thenavigation source is an SRN. For an invalid course select, three amberdashes (---) replace the readout digits.

(b) Desired Track Readout (FMS)

The desired track digital readout has a range from 001° to 360° with aresolution of one degree. The readout digits are displayed in magenta. Amagenta DTK label annunciates the digital readout. For an invalid desiredtrack, three amber dashes (---) replace the readout digits.

(c) Course Select/Desired Track Pointer (VOR/ILS/FMS)

The pilot or copilot uses the course knob on the GC-550 Guidance ControlUnit to position the pointer to a magnetic bearing that coincides with thedesired VOR radial or localizer course. The pointer rotates around thestationary aircraft symbol to provide a continuous readout of course error tothe IAC flight director. The pointer is displayed in the same color as thecourse/desired track readout.

The PFD displays the entire pointer when the full heading compass display isselected. Portions of the pointer may be out of view when the partial headingcompass display is selected. The pointer slews to the current navigationsource bearing when the course knob on the GC-550 Guidance Control Unit ispushed and the navigation source is an SRN.



The PFD removes the pointer for the following conditions:

• The selected navigation source is an SRN and course select is invalid

• The selected navigation source is an FMS and the desired track is invalid.

(5) Drift Bug

The PFD positions the magenta drift bug on the outer edge of the compass arc. Thebug moves around the compass perimeter to represent the angular differencebetween FMS computed track and actual aircraft track. Increasing values of driftcause a clockwise rotation of the bug. The range of the drift bug is as follows:

• The drift bug has a display range of ±180° for the full heading compass display

• The drift bug has a display range of ±45° for the partial heading compassdisplay.

The PFD removes the drift bug for an invalid drift angle or an invalid heading.

(6) Bearing Pointers and Source Identifiers

The PFD presents bearing information as two independent pointers on the HSIdisplay. The bearing source select knobs on the on-side DC-550 Display Controllercontrol the position of the pointers. The VOR and FMS bearing pointers are cardreferenced. The ADF bearing pointers are case referenced. Increasing values ofbearing cause clockwise movement of the bearing pointers. The bearing select datadefaults to VOR1 on the left side and VOR2 on the right side when the displaycontroller is invalid.

(a) Bearing Pointer No. 1

The PFD displays bearing pointer No. 1 as two single, cyan bars separated bythe length of the lateral deviation bar, and extending through the center of theaircraft symbol. A circle symbol is drawn below the head of the pointer. Thebearing pointer points to the current bearing position. The bearing pointerrotates around the center of the aircraft symbol with 360 degrees of motion.

The entire bearing pointer is in view on the full heading compass display. Onthe partial heading compass display, only one of the two bars is visible at agiven time. The PFD removes the pointer from the display if the bearingsource select knob on the DC-550 Display Controller is set to off. The PFDalso removes the pointer from the display if the bearing source is VOR and theradio frequency is TTL.

For an invalid bearing, the PFD removes the bearing pointer from the display.If the pilot’s display controller is invalid, then the bearing pointer on the pilot’sPFD defaults to VOR1. If the copilot’s display controller is invalid, then thebearing pointer is removed from the copilot’s PFD.



(b) Bearing Pointer No. 2

The PFD displays bearing pointer No. 2 as a double set of white parallel bars,separated vertically by the length of the lateral deviation bar and laterally bythe width of the diamond symbol. The diamond symbol is drawn below thehead of the pointer. A long tick mark extending from the diamond points to thecurrent bearing position. A T-shaped tick mark at the base of the parallel barspoints to the reciprocal of the current bearing. The bearing pointer rotatesaround the center of the aircraft symbol with 360 degrees of motion.

The entire bearing pointer is in view on the full heading compass display. Onthe partial heading compass display, only one of the two sets of bars is visibleat a given time. The PFD removes the pointer from the display if the bearingsource select knob on the DC-550 Display Controller is set to off. The PFDalso removes the pointer from the display if the bearing source is VOR and theradio frequency is TTL.

For an invalid bearing, the PFD removes the bearing pointer from the display.If the copilot’s display controller is invalid, then the bearing pointer on thecopilot’s PFD defaults to VOR2. If the pilot’s display controller is invalid, thenthe bearing pointer is removed from the pilot’s PFD.

(c) Bearing No. 1 Source Identifier

The PFD display a circle to the left of the bearing source No. 1 identifier field.The identifier field displays the currently selected bearing source for bearingpointer No. 1. The circle and bearing source identifier are displayed in cyan.Bearing source identifiers are listed in Table 2-1-27. The PFD removes thebearing source identifier and circle if the bearing source select knob on theDC-550 Display Controller is set to off.

Table 2-1-27. Bearing Source No. 1 Identifier

Bearing Source KnobSelection

Bearing SourceIdentifier

Aircraft Reference

OFF No Identifier N/A

NAV1 VOR1 Card

ADF ADF1* Case

* ADF if single installation



(d) Bearing No. 2 Source Identifier

The PFD displays a diamond the left of the bearing source No. 2 identifierfield. The identifier field displays the currently selected bearing source forbearing pointer No. 2. The diamond and bearing source identifier aredisplayed in white. Bearing source identifiers are listed in Table 2-1-28. ThePFD removes the bearing source identifier and diamond if the bearing sourceselect knob on the DC-550 Display Controller is set to off.

Table 2-1-28. Bearing Source No. 2 Identifier

Bearing Source KnobSelection

Bearing SourceIdentifier

Aircraft Reference

OFF No Identifier N/A

NAV2 VOR2 Card

ADF ADF2* Case

* ADF if single installation

(7) Distance Readout

The PFD displays the distance readout to the left of the heading compass display.The readout displays the distance to a navigation source station for SRN sources, orthe distance to a waypoint for FMS sources. Distances are shown in nautical milesalong with a white NM annunciation. The PFD replaces the NM annunciation with anamber H annunciation when DME hold is active. The DME hold mode is selectedthrough the radio management unit. For an invalid DME hold annunciation, the PFDremoves the H annunciation from the display.

(a) Distance Readout (SRN)

The distance readout range is from 0.0 to 399.9 NM for an SRN navigationsource. Resolution is in 0.1 NM increments from 0 to 100 NM, and in 1.0 NMincrements for distances greater than 100 NM. The readout digits aredisplayed in green. For invalid DME distances, three amber dashes (---)replace the readout digits.

(b) Distance Readout (FMS)

The distance readout range is from 0.0 to 4095 NM. Resolution is in 0.1 NMincrements for distance values less than 100 NM, and in 1 NM increments fordistance values greater than 100 NM. The readout digits are displayed inmagenta. for invalid FMS distances, three amber dashes (---) replace thereadout digits.



(8) Distance Identifier (FMS)

When FMS is the selected navigation source, the PFD displays a magenta distanceidentifier that represents the TO waypoint. The PFD removes an invalid TO waypointidentifier from the display.

(9) Time-To-Go Readout

The PFD displays the time-to-go (TTG) in a digital readout that shares the samelocation as the ground speed and elapsed time digital readouts. Time is displayed inminutes rounded to the nearest minute. The readout range is from 0 to 399 minutesif the selected navigation sources is an SRN, and 0 to 512 minutes if the selectednavigation sources is an FMS. The readout resolution is one minute. A white TTGlabel annunciates the readout.

The readout digits are displayed in green if the on-side SRN is the navigationsource, and in yellow if the cross-side SRN is the navigation source. The readoutdigits are displayed in magenta if the FMS is the navigation source. A white MINlabel is displayed next to the readout digits. For invalid time-to-go data, three amberdashes (---) replace the readout digits.

(10) Ground Speed Readout

The PFD displays the ground speed (GSDP) in a digital readout that shares thesame location as the time-to-go and elapsed time digital readouts.. Ground speed isdisplayed in knots rounded to the nearest knot. The readout range is 0 to 999 knotsif the selected navigation sources is an SRN, and 0 to 4095 knots if the selectednavigation source is an FMS. The display resolution is one knot. A white GSPDlabel annunciates the readout.

The readout digits are displayed in green if the on-side SRN is the navigationsource, and in yellow if the cross-side SRN is the navigation source. The readoutdigits are displayed in magenta if the FMS is the navigation source. A white KTSlabel is displayed next to the readout digits. For an invalid ground speed, threeamber dashes (---) replace the readout digits.

(11) Elapsed Time Readout

Elapsed time is computed using an internal time base. The PFD displays elapsedtime in a digital readout that shares the same location as the time-to-go and groundspeed readouts. The readout range is from 0:0 (minutes:seconds) to 9:59(hours:minutes). Leading zeros are displayed for values less than 10. A white ETlabel annunciates the readout.

The ET pushbutton on the DC-550 Display Controller allows the pilot or copilot todisplay the elapsed time. The ET pushbutton is also used to reset the elapsed timereadout. When elapsed time is less than one hour, the readout format is MM:SS,where MM represents elapsed minutes and SS represents elapsed seconds. Whenelapsed time is greater than one hour, the readout format is HH:MM, where HHrepresents elapsed hours and MM represents elapsed minutes. The PFD removesthe elapsed time readout if the display controller is invalid.



(12) Wind Display

The FMS provides the wind display information. The PFD then displays the windspeed in a digital readout and the wind angle as an arrow. The readout and arroware displayed to the left of the heading compass scale, just below the distancereadout. The wind display is provided for both SRN and LRN modes.

(a) Wind Speed Readout

The PFD positions the wind speed readout to the left of the wind angle arrow.Wind speed is displayed in knots. The readout has a range from one knot to255 knots with a resolution of one knot. The readout digits are displayed inmagenta. The PFD removes the wind speed readout for an invalid windspeed, or when the wind speed readout indicates 0 knots.

(b) Wind Angle Arrow

The PFD positions the magenta wind angle arrow to the right of the windspeed readout. The wind angle is given in degrees. An increasing wind anglecauses a clockwise rotation of the arrow. The arrow rotates about its centerwith 360 degrees of motion. The wind angle arrow is displayed in magenta.The PFD removes the wind angle arrow for an invalid wind speed, or when thewind speed readout indicates 0 knots.

(13) Weather Radar Display

The pilot or copilot uses the FULL/WX pushbutton on the DC-550 Display Controllerto select the weather radar display for the PFD. The weather radar (WX) display iscase-referenced. The WX receiver/transmitter provides information on two controlbuses, a serial control interface (SCI) to the IAC and a WX video interface bus to thePFD. The SCI bus contains the WX modes and control information needed toprovide status indications about the weather radar receiver/transmitter on the PFD.The WX video interface bus provides the WX video data for display on the PFD.Figure 2-1-18 shows the partial compass display with weather radar data.

(a) WX Half-Range Ring

When the WX video data is selected for display, the PFD draws a white ±45degree arc on the partial heading compass display. This arc represents theWX half-range ring, and is positioned halfway between the aircraft symbol andthe compass card boundary. The digital half-range value is centered belowthe right edge of the half-range ring.

The pilot or copilot selects the weather radar range using the control knob onthe weather radar controller. The weather radar range is limited to 5 NMthrough 1,000 NM. The weather radar range defaults to 50 NM for an invalidweather radar control bus. If the weather radar range is invalid, it defaults to100 NM (50 NM on the half-range ring). Table 2-1-29 lists the availableweather radar ranges that can be selected.



Table 2-1-29. Weather Radar Ranges

Weather Radar Selection(NM)

Half-Range Ring Displayed Range(NM)

5.0 2.5

10.0 5

25.0 12.5

50.0 25

100.0 50

200.0 100

300.0 150

500.0* 250

1,000.0* 500

* Flight plan mode on WX controller selected

(b) WX Video Data

The PFD displays the WX video data in a 120° pattern if sector scan has notbeen selected on the weather radar controller. If sector scan has beenselected, the PFD displays WX video data in a 60° pattern. The 60° scan isfurther identified by two white azimuth marks on the half-range ring at ±30degrees to either side of an imaginary line running through the center of theaircraft symbol. Table 2-1-30 specifies the display colors for weather radarand ground map returns.

Table 2-1-30. Color Codes for Weather Radar Data

Return WX Mode GMAP Mode

Level 0 Black Black

Level 1 Green Cyan

Level 2 Yellow Yellow

Level 3 Red Magenta

Level 4 Magenta Black (N/A)

REACT Cyan Black (N/A)

Turbulence White Black (N/A)




.410 M 29.92 IN

DTK359

HDG001

0

1

23

32

133N

3603000

1000

VOR1ADF2

12.5H

25

GSPD245KTS

MAG2

260

235

245

200

260

280

220

YDIASHDG 14500

FMS

20

20 20

20

10

10 10

10

2401

9

GSAP

14500

0020

80 143

AD-51193@

WX VIDEODISPLAY

LOC

TGT

DR

KDVT

WX3.5

50

FMS

WXHALF-RANGE

RING

WXTILT ANGLEREADOUT

WXWARNING

ANNUNCIATOR

WX MODEANNUNCIATOR

350

Figure 2-1-18. Partial Heading Compass Display With Weather Radar Data



A weather radar fault by itself does not remove or prevent weather radarreturns. If the PFD is not receiving weather radar data, it erases the weatherradar display. The PFD also removes the weather radar display for an invalidweather radar control bus or an invalid weather radar range. The PFD clearsthe weather radar display for range changes, transitions into 60° scan, and ontransitions into or out of a system test or the GMAP mode.

(c) WX Mode Annunciator

The WX mode annunciator is centered below the left edge of the half-rangering. The PFD displays WX mode annunciations only when the weather radardisplay is selected. Table 2-1-31 lists the WX mode annunciations and theircorresponding display colors.

Table 2-1-31. PFD WX Mode Annunciations

Annunciation Color Mode Description

CR/R Green Normal weather radar mode with GCR andRCT

FAIL Amber Test mode and faults

FPLN Green Flight plan mode

FSBY Green Forced standby mode (WOW)

GCR Green Normal weather radar mode with groundclutter rejection (GCR)

GMAP Green Ground map mode

RCT Green Normal weather radar mode with REACT

R/T Green Weather radar with REACT and turbulence

STBY Green Normal standby mode

TEST Green Test mode and no faults

WAIT Green Power-up mode for approximately oneminute

WX Green Normal weather radar mode

WX Amber Invalid weather radar or invalid weatherradar control bus

WX/T Green Normal weather radar with turbulence



(d) WX Tilt Angle Readout

The PFD displays the WX tilt angle readout only when the weather radardisplay is selected. The readout range is -16° to +15° with a resolution of onedegree. The PFD displays a plus (+) sign in front of positive values, and aminus (-) sign in front of negative values. A degree symbol (°) is displayedwith the tilt angle value if auto-tilt is not active. If auto-tilt (a WU-870receiver/transmitter feature) is active, an A is displayed in place of the degreesymbol. The readout digits are displayed in green.

A WX fault in a non-test mode does not cause the PFD to remove or disablethe tilt angle readout. The PFD removes the tilt angle readout when the WXtest mode is selected and faults are detected. The PFD also removes the tiltangle readout for an invalid weather radar control bus or an invalid tilt angle.

(e) WX Warning Annunciator

The PFD displays WX warning annunciations if the standby modes are notselected and the appropriate warning is selected via the weather radarcontroller. The target alert annunciation has priority over the variable gainannunciation. For an invalid weather radar control bus, the PFD removes anycurrently displayed WX warning annunciation. Table 2-1-32 lists the WXwarning annunciations.

Table 2-1-32. WX Warning Annunciations

Annunciation Color Description

VAR Amber Variable Gain

TGT Green Target Alert Enable

TGT Amber Target Alert enabled and Level 3WX return detected in forward 15°of antenna scan

(f) WX Fault Codes

The PFD removes the WX tilt angle when the WX test mode is selected andfaults are detected. Instead, the PFD displays fault codes in the WX tilt anglereadout area. The WX receiver/transmitter sends the fault codes to the PFDfor display. The fault codes are displayed in amber.



E. PFD Air Data Display Operation

The PFD displays air data information in three prime viewing areas. Air data informationincludes barometric altitude, indicated airspeed, and vertical speed.

(1) Altitude Display

The PFD presents the aircraft’s barometric altitude indication as a vertical altitudetape display. The altitude tape display has chevrons positioned at 500 feet and1,000 feet increments to indicate the climb or descent rate to the pilot or copilot.The altitude tape display also has a rolling digital readout for accurate interpretationof the barometric altitude. In addition, the altitude display has an altitude trendvector, an altitude preselect readout, an altitude preselect bug, and a barometriccorrection readout. Figure 2-1-19 shows the PFD altitude display format.

The AZ-850 MADC provides the inputs for the barometric altitude, altitude rate, andbarometric correction. The flight director function provides the input for altitudeselect.

(a) Altitude Scale and Altitude Tape

The altitude scale and altitude tape provide a trend indication of the aircraft’scurrent barometric altitude. The PFD displays the altitude tape as avertically-oriented rectangle with gray shading. White tick marks representingthe scale are displayed at 100-feet increments on the inside-left edge of thealtitude tape.

White, single line chevrons are present within the altitude tape in 500 footincrements. White, double line chevrons are present within the altitude tape in1,000 foot increments. A vertical line connects the chevrons. White digits arecentered and left-justified within the chevrons at 500 foot increments.

A white box outlining the rolling altitude digital readout is centered within thealtitude tape. The altitude tape moves vertically behind the digital readout,and displays barometric altitude ±550 feet from the current altitude. Thealtitude tape can display altitudes from -1,000 feet to +60,000 feet, with largernumbers displayed at the top of the tape. Increasing values of barometricaltitude cause downward movement of the altitude tape.

For invalid barometric altitudes, the PFD removes the readout digits, and thetick marks and chevrons on the altitude tape. The PFD also draws a red Xover the altitude tape and scale.



(b) Altitude Rolling Digital Readout

A rolling digital readout of current barometric altitude is centered within thealtitude tape. The readout can display altitudes from -1,000 to +60,000 feet.Display resolution is 20 feet with additional resolution provided by rolling thedigits. Altitude values between 0 and 10,000 feet have a hashed box in theten-thousands position. This box serves as a place holder. Negative numbershave a right-justified minus sign (-). The readout digits are displayed in green.

29.92 IN

14500

14000

14500

0020

80143

ALTITUDE SELECTDIGITAL READOUT

ALTITUDE SCALEAND TAPE

500 FOOTCHEVRON

ALTITUDESELECT BUG

ALTITUDE ROLLINGDIGITAL READOUT

1000 FOOTCHEVRON

BARO SETDIGITAL READOUT

ALTITUDETREND VECTOR

AD-51194@NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS

Figure 2-1-19. PFD Altitude Display

(c) Altitude Trend Vector

The magenta altitude trend vector provides an indication of the altitude rate.The input is scaled so that the altitude trend vector represents the altitude thatthe aircraft should attain in six seconds if the current altitude rate (verticalspeed) is maintained. The PFD displays the altitude trend vector at thevertical center of the altitude tape, outside the left edge of the altitude tapeoutline. For an invalid altitude rate, the PFD removes the altitude trend vectorfrom the display.

(d) Altitude Select Digital Readout

The PFD displays the altitude select digital readout at the top of the altitudetape. The readout has a range from -900 to +51,000 feet, with a displayresolution in 100 feet increments. A white box outlines the readout digitswhich are displayed in cyan for normal conditions. The box and readout digitschange to amber for alert conditions. For an invalid altitude select, five amberdashes (-----) replace the readout digits.



(e) Altitude Select Bug

The PFD displays the cyan altitude select bug on the left edge of the altitudetape. The bug moves vertically along the tape in 100 foot increments. Thebug is limited to 550 feet when the altitude tape displays a barometric altitude±550 feet from the current altitude. The portion of the bug past the 550 footlimit is masked. For an invalid altitude select, the PFD removes the bug fromthe display.

(f) Altitude Alert

The master flight director controls the altitude alert annunciation. Alertannunciations consist of an audio output, visual indications, and a colorchange to the altitude select digital readout. As the aircraft approaches within1,000 feet of the selected altitude, the readout changes colors until the aircraftis within 250 feet of the selected altitude. The outline box around the readoutand the readout digits change colors as follows to provide an altitude alertwarning to the pilot or copilot.

Color Alert State

White (box)Green (digits)

Normal

Amber Alert

(g) Baro Set Digital Readout

The PFD displays the baro set digital readout at the bottom of the altitudetape. The readout can display both inches of mercury and HectoPascals. Thepilot or copilot uses the BARO knob on the PFD bezel controller to set thebarometric correction value in the readout. Pushing the IN/HPA button on thebezel controller toggles the readout between inches of mercury andHectoPascals. Table 2-1-33 specifies the range and resolution of the digitalreadout.

The readout digits are displayed in cyan along with a white IN or HP label toindicate inches of mercury (IN) or HectoPascals (HP). Increasing values ofbarometric correction cause a decrease in the barometric altitude readout. Foran invalid barometric correction, four amber dashes (--.--) replace the readoutdigits if inches of mercury are selected, or three amber dashes (---) replacethe digits if HectoPascals are selected.

Table 2-1-33. Barometric Correction Range and Resolution

Selected Units Range Resolution Units Label

Inches of Mercury 16.00 to 32.00 0.01 IN

HectoPascals 541 to 1083 1 HPA



(2) Airspeed Display

The PFD presents the aircraft’s airspeed indication as a vertical tape display. Thevertical tape display provides a scrolling tape and digital readout for accurateinterpretation of airspeed. In addition, the airspeed display has an airspeed trendvector, a selected airspeed digital readout and bug, overspeed indications, Vspeedbugs, an IAS/Mach digital readout, and Vspeed digital readouts. Figure 2-1-20shows the PFD airspeed display format.

The AZ-850 MADC provides the inputs for calibrated airspeed, Mach, Vmo, andMmo. The flight director function provides the input for the selected airspeed. Theinputs for V1 (critical engine failure speed), V2 (takeoff climb speed), and VR(rotation speed) are derived from the MFD bezel controller through the DC-550Display Controller. The stall warning computer provides an angle of attack(normalized) input through the DA-800 DAU.

(a) Airspeed Scale and Airspeed Tape

The airspeed scale and tape provide a trend indication of the aircraft’sindicated airspeed (IAS). The altitude tape is displayed as a vertically-orientedrectangle with gray shading. White tick marks representing the scale areshown at 10-knot increments on the inside-right edge of the airspeed tape.Digits are right-justified next to the tick marks at 20-knot increments, startingat 40 knots.

A T-shaped box, centered within the airspeed tape, surrounds the rollingdigital readout of current airspeed. The airspeed tape moves vertically behindthe rolling digital readout, and displays airspeed ±42 knots from the currentairspeed. The airspeed tape displays airspeeds from 40 knots to 450 knots,with larger numbers displayed at the top of the tape. The airspeed tape parksat 40 knots and 450 knots for airspeeds below 40 knots and above 450 knots,respectively.

For an invalid indicated airspeed, the PFD removes the readout digits, and thetick marks and chevrons on the airspeed tape. The PFD also draws a red Xover the airspeed tape and scale.

(b) Airspeed Rolling Digital Readout

A rolling digital readout of airspeed is located within the T-shaped boxcentered within the tape proper. The digits are shown with the ones positionrepresented by rolling digits. The readout displays airspeeds from 40 to 400knots, with a resolution of one knot. Additional resolution is provided by rollingthe ones digits. Increasing values in the readout indicate a greater airspeedand cause the ones digit to roll downward.

The readout digits are displayed in green for normal airspeeds. The readoutdigits are displayed in amber if the trend vector exceeds Vmo. The readoutdigits are displayed in red if the airspeed is equal or greater than Vmo.



R

160

35

55

60

80

135

401

9R1

2

.410 M

260

235

245

200

260

280

220

2401

9

AOA

R1

2

.410 M

260

235

245

200

260

280

220

2401

9

2

125

1120

VMO/MMOINDICATOR

LOW AIRSPEEDAWARENESS

INDICATOR

AOAANNUNCIATOR

AIRSPEEDSCALE AND

TAPE

AIRSPEEDROLLINGDIGITAL

READOUT

VSPEEDDIGITAL

READOUTS

AIRSPEEDREFERENCE

DIGITALREADOUT

ACCELERATIONTREND VECTOR

V2 SET BUG

VR SET BUG

V1 SET BUG

MACHDIGITAL

READOUT

AIRSPEEDREFERENCE

BUG

AD-51195@

.410 M

AP SET BUG

L

Figure 2-1-20. PFD Airspeed Display



(c) Airspeed Reference Digital Readout

The PFD displays the airspeed reference in a digital readout above theairspeed tape. A white box outlines the readout. The readout, along with theairspeed reference bug, provides a full time display for the pilot independent ofthe flight director mode. However, a selected airspeed reference can only beset when the speed mode is engaged in either flight director. The readoutrange is 80 knots to Vmo knots with a resolution of 1 knot when IAS is theselected speed mode. The readout range is 0.20 Mach to Mmo with aresolution of 0.01 Mach when Mach is the selected speed mode.

The SPD control knob on the GC-550 Guidance Control Unit controls theairspeed reference digital readouts and reference bugs on each PFD.Clockwise knob rotation changes the readout digits in one knot increments forIAS, or 0.1 increments for Mach. Counter-clockwise knob rotation changes thereadout digits in one knot decrements for IAS, or 0.1 decrements for Mach.The knob is also an integral pushbutton (PUSH IAS/M) which toggles thereadout between IAS and Mach when pushed. A cyan M follows the readoutdigits when Mach is the selected speed mode.

The readout digits are displayed in cyan for normal conditions. If the aircraftspeed exceeds Vmo, the readout digits are displayed in red. If the airspeedtrend vector exceeds Vmo, the readout digits are displayed in amber. For aninvalid selected airspeed, three amber dashes (---) replace the readout digitswhen the selected speed mode is IAS. When the selected speed mode isMach, two amber dashes (--M) replace the readout digits for an invalidselected Mach.

(d) Airspeed Reference Bug

The PFD positions the cyan airspeed reference bug on the right edge of theairspeed tape. The bug moves vertically along the airspeed tape with eachone knot increment in IAS or 0.01 increment in Mach as selected airspeedchanges. The bug is limited to 42 knots or 0.107 Mach from the currentlydisplayed airspeed value. If the selected airspeed reference is set to a valueoutside the display range of the airspeed tape, the bug parks itself at the endof the tape, half visible. The PFD removes the bug for an invalid selectedairspeed, regardless of whether IAS or Mach is the selected speed mode.

(e) Acceleration Trend Vector

The magenta acceleration trend vector provides an indication of theacceleration direction. The input is scaled so that the trend vector representsthe airspeed that the aircraft would attain in 10 seconds if the current aircraftacceleration is maintained. The trend vector is displayed at the vertical centerof the airspeed tape, outside the right edge of the airspeed tape outline. Themaximum movement of the trend vector is 42 knots from the present airspeed.The trend vector is inhibited on the ground during the takeoff phase. For aninvalid airspeed, the PFD removes the trend vector from the display format.



(f) Vmo/Mmo Indicator

The Vmo/Mmo indicator provides an indication of overspeed conditions. ThePFD displays the Vmo/Mmo indicator as a red bar inside the airspeed tape.The bar extends from the Vmo/Mmo tape position to the top of the tape. If theaircraft speed exceeds Vmo or Mmo, the digits in the airspeed referencedigital readout and Mach digital readout are displayed in red. If the airspeedtrend vector exceeds Vmo or Mmo, amber digits are displayed in bothreadouts. For an invalid airspeed or Vmo, the PFD removes the Vmo/Mmoindicator from the display format.

(g) Low Airspeed Awareness Indicator

The low speed awareness represents the aircraft stall speed. The stall speedis calculated as a function of the angle of attack (AOA) provided by the stallwarning computer via the DAU. The PFD displays the low airspeed awarenessindicator as a red, amber, and white thermometer located inside the airspeedtape, against the right-hand edge. The white band (top) extends from 1.13 to1.23 times Vstall. The amber band (middle) extends from 1.0 to 1.13 timesVstall. The red band (bottom) extends from 1.0 times Vstall to the bottom ofthe airspeed tape.

The PFD removes the low speed awareness indicator from the display whenweight on wheels indicates an on ground status. The PFD also removes thelow speed awareness indicator for an invalid indicated airspeed or angle ofattack. For an invalid angle of attack or indicated airspeed, the PFD displaysan amber AOA annunciation next to the airspeed tape.

(h) Vspeed Set Bugs

Three Vspeed set bugs can be displayed on the airspeed scale and setthrough the MFD bezel controller. These bugs allow the pilot or copilot tovisually monitor key airspeed references (Vspeeds), which are further definedas V1, VR, and V2. The magenta V1 bug is the decision speed reference bug.The cyan VR bug is the rotation speed reference bug. The white V2 bug isthe safety speed reference bug.

The Vspeed bugs move vertically with respect to the airspeed tape. When aVspeed bug set value is within 42 knots of the current IAS, the bug isdisplayed at the airspeed scale position on the perimeter of the airspeed tape.The bug is not visible if the Vspeed bug set value is not within 42 knots.

The PFD removes the Vspeed bugs from the display if they are currently beingdisplayed and the airspeed increases beyond 230 knots. The PFD alsoremoves the Vspeed bugs from the display for an invalid airspeed or invalidVspeeds.



(i) Vspeed Digital Readouts

Along with the Vspeed bugs, the PFD displays a digital readout of eachreference speed in the lower half of the airspeed tape. The readout digits aredisplayed in the same color as the corresponding Vspeed bug which ispositioned to the right of the readout. The PFD positions the readouts inascending order (V1, VR, V2), starting at the bottom of the airspeed tape.

On power-up, the digital readout displays three amber dashes (---). Thedashes change to digits when the respective Vspeed is set by the MFD bezelcontroller. The PFD removes the Vspeed readouts from the format for aninvalid airspeed. The PFD removes the readouts from the display when theairspeed increases to a value where the first set bug is visible on the airspeedtape. For invalid Vspeeds, three amber dashes (---) replace the readoutsdigits.

(j) Mach Digital Readout

The PFD displays a digital readout of Mach below the airspeed tape. Thereadout has a range from 0.050 to 1.000 Mach with a resolution of 0.001Mach. A white M is displayed to the right of the readout value. Hysteresis isapplied to the Mach value such that the Mach readout is displayed when anincreasing Mach number exceeds 0.45. The Mach readout remains displayeduntil the Mach value drops below 0.05, at which time the readout is removed.

The readout digits are displayed in green for normal conditions. If the aircraftspeed exceeds Mmo, the readout digits are displayed in red. If the airspeedtrend vector exceeds Mmo, the readout digits are displayed in amber. For aninvalid selected Mach, three amber dashes (---) replace the readout digitswhen the selected speed mode is Mach.

(3) Vertical Speed Display

The PFD presents the aircraft’s vertical speed as an analog arc display. The arcdisplay includes a vertical speed scale with a dynamic pointer, a digital readout, anda vertical speed reference digital readout and bug. The MADC provides the input forthe altitude rate. The flight director function provides the input for the vertical speedreference data. Increasing positive values of vertical speed indicate a climb. Figure2-1-21 shows the PFD vertical speed display.

(a) Vertical Speed Scale and Pointer

The vertical speed scale and pointer provide an analog indication of theaircraft’s vertical speed. The scale is displayed as a white 134° arc with digitsrepresenting thousands of feet per minute of vertical speed displayed at thefollowing scale positions: 0, ±1,000, ±2,000, and ±3,000. Only the thousandsposition of each digit is displayed. The scale is non-linear to provideincreased resolution around zero vertical speed.



Tick marks representing 500 feet per minute increments are displayed on thearc from -3,000 to +3,000 feet per minute. A small tick mark identifies the 500feet per minute positions. A large tick mark identifies the ±1,000, ±2,000, and±3,000 feet per minute positions.

A pointer is continuously displayed on the vertical speed scale to indicate thevertical speed of the aircraft. Pointer movement is non-linear; greaterresolution is provided for vertical speeds between ±1,000 feet/minute. Displayresolution is highest between ±1,000 feet/minute, and lowest outside of ±3,000feet/minute. For vertical speeds less than -8,100 feet/minute or greater than+8,100 feet/minute, the pointer is pegged at the corresponding negative orpositive 8,100 feet/minute position (endpoints of the scale). The pointer isdisplayed in green for normal vertical speed conditions, and in red for alertconditions.

For an invalid vertical speed, the PFD displays the vertical speed scale asfollows:

• The pointer is removed from the display

• A box is displayed vertically in the center of the scale with red letters VScentered in the box.

0

1

23

32

1

3000

1000

VERTICAL SPEEDREFERENCEDIGITAL READOUT

VERTICAL SPEEDREFERENCE BUG

VERTICAL SPEEDSCALE AND POINTER

VERTICAL SPEEDDIGITAL READOUT

AD-51196@

Figure 2-1-21. PFD Vertical Speed Display



(b) Vertical Speed Digital Readout

The PFD displays a digital readout of the actual vertical speed in the center ofthe vertical speed scale. The readout range is from -9,999 to +9,999feet/minute with a resolution of 50 feet/minute. The PFD removes the readoutfrom the display for vertical speeds between -550 feet/minute and +550feet/minute. The readout digits are displayed in green for normal verticalspeed conditions, and in red for alert conditions.

(c) Vertical Speed Reference Digital Readout

The PFD displays the vertical speed reference digital readout above thevertical speed scale when the vertical speed mode is engaged in the flightdirector. A white box outlines the readout. The readout has a range from-6,000 feet per minute to +6,000 feet per minute, with a display resolution of100 feet per minute. The readout digits are displayed in cyan.

The pilot or copilot uses the SPD knob on the GC-550 Guidance Control Unitto select a vertical speed reference. A master/slave condition exists such thatthe master IAC drives both the pilot’s and copilot’s vertical speed referencereadout. If neither flight director is engaged in the vertical speed mode, thePFD removes the vertical speed reference readout and reference bug fromboth PFDs. The PFD also removes the vertical speed reference digitalreadout for an invalid vertical speed reference.

(d) Vertical Speed Reference Bug

The PFD displays the cyan vertical speed reference bug when the verticalspeed mode is engaged in the flight director. The bug is positioned along theinside edge of the vertical speed scale. The bug travels from -6,000 feet perminute to +6,000 feet per minute in 100 feet per minute increments. For aninvalid vertical speed reference, the PFD removes the reference bug.



F. PFD TCAS Display Operation

This section defines the symbology (resolution advisories, traffic, and annunciations) that isused on the PFD for the Traffic Alert and Collision Avoidance System (TCAS). Data isreceived from the TCAS computer via a high speed ARINC 429 input bus. Figure 2-1-22shows the TCAS symbology and annunciations that are displayed on the PFD.

The PFD displays TCAS resolution advisories (RA) and general TCAS operating modesand failure annunciations. An RA is a display indication given to the pilot recommending orprohibiting a maneuver to hazardously close encounters with an intruding aircraft. Thereare two types of RAs: corrective and preventative. A corrective RA instructs the pilot todeviate from current vertical speed to avoid the intruder. A preventative RA instructs thepilot to avoid certain deviations from the current vertical speed. The TCAS computerprovides RA directions in the form of vertical speed commands, which are displayed in thevertical speed scale.

(1) TCAS Vertical Speed Indications (VSI)

Resolution advisories consists of one or two red bands, and up to one green bandlocated on the inside of the vertical speed scale. The red bands signify verticalspeeds that should be avoided if the RA is preventive, or regions that should beimmediately flown from if the RA is corrective. The green band provides a "fly-to"indication which is a vertical speed command cue for the pilot during corrective RAs.The colors of the vertical pointer and digital readout reflect the color of the TCAS RAband that the pointer is currently in.

(a) Red Band(s)

The PFD places the red band(s) on the inside of the vertical speed scale asshown in Figure 2-1-22. The red bands are comprised of two thick lineswhose length are determined by information in the vertical RA input data. Thedown advisory red band starts at the upper display limit of the vertical speedscale and extends down to the specified value. A red tick mark is displayed atthe upper display limit. The up advisory red band starts at the lower displaylimit of the vertical speed scale and extends up to the specified value. A redtick mark is displayed at the lower display limit.

The red bands are only displayed for each valid up or down advisory asdefined in Table 2-1-34. For the case where the up advisory states don’tclimb and the down advisory states don’t descend, the red stroked zone from-8,100 to -250 feet/minute and from +250 to +8,100 feet/minute is displayed.A red tick mark is displayed at the most positive and negative ends of thescale. In the preventative case, the remainder of the vertical speed scale thatis not a red band is displayed as a white band. The PFD removes the redband(s) from the display format for an invalid VSI TCAS display.




260

R

TCASTEST

14000

ILS1

VERTICAL SPEEDDIGITAL READOUT

TCAS MODEANNUNCIATOR

235

245

359

E

W

21

12

30

15

33

S

24

200

260

280

220 1

2

.410 M

VOR1

ADF2

YDGSLOC 14500

14500

GS

3

1000

1

2

0

1

2

3

0020

80240

1

9

AD-51197@

CAT2AP

143

N3

6

CRS

TTG5MIN

20

10

10

20

40

GREEN BAND(TCAS "FLY-TO"ZONE BAND)

RED BANDS(TCAS RA BANDS)

Figure 2-1-22. PFD TCAS Display



Table 2-1-34. Resolution Advisory Matrix

TCAS RADown

Advisory

TCAS RA Up Advisory

TCAS RACombined

ControlNo Up

AdvisoryClimb Don’t

DescendDon’t

Descend>500

Don’tDescend

>1,000

Don’tDescend

>2,000

No DownAdvisory

NVNVNV

VNVV

VNVV

VNVV

VNVV

VNVV

CLBCorrective

DESCCorrective

Preventative

Descend NVVV

NVNVNV

NVNVNV

VVV

VVV

VVV

CLBCorrective

DESCCorrective

Preventative

Don’tClimb

NVVV

NVNVNV

VVV

VVV

VVV

VVV

CLBCorrective

DESCCorrective

Preventative

Don’tClimb>500

NVVV

VVV

VVV

VVV

VVV

VVV

CLBCorrective

DESCCorrective

Preventative

Don’tClimb

>1,000

NVVV

VVV

VVV

VVV

VVV

VVV

CLBCorrective

DESCCorrective

Preventative

Don’tClimb

>2,000

NVVV

VVV

VVV

VVV

VVV

VVV

CLBCorrective

DESCCorrective

Preventative

CLB = Climb DESC = Descend V = Valid NV = Not Valid



(b) Green Band

When the TCAS RA combined control indicates a climb corrective RA, thegreen "fly-to" zone is attached to the end of the red zone defined by the TCASRA up advisory. When the TCAS RA combined control indicates a descendcorrective RA, the green "fly-to" zone is attached to the end of the red zonedefined by the TCAS RA down advisory. The green "fly-to" zone is notdisplayed when the TCAS RA combined control indicates a preventative RA.The green band is four lines wide and equals the linear distance from the1,500 to 2,000 feet/minute distance on the vertical speed scale.

A special case exists in which the green band is placed at a fixed position onthe scale. This occurs when the TCAS RA down advisory states don’t climband the TCAS RA up advisory states don’t descend, which may happen whileflying around the zero feet/minute vertical rate. A green band is forced ontothe vertical speed scale, centered at zero feet/minute, to provide a clear safeband for the pilot. The pilot then keeps the current vertical speed in this safeband zone. The red bands have endpoints at the edges of the green band(approximately ±250 feet/minute) in this case. The PFD removes the greenband from the display format for an invalid VSI TCAS display.

(2) TCAS Mode Annunciator

The PFD displays TCAS mode annunciations based on conditions listed in Table2-1-35. The TCAS mode annunciations are displayed with the priority listed in thetable: highest first and lowest last.

Table 2-1-35. PFD TCAS Mode Annunciations

Annunciation Color Condition

TCAS FAIL Amber Indicates that TCAS data is invalid.

TCAS TEST White Indicates that the TCAS is undergoing a functionaltest.

TCAS OFF White Indicates that the TCAS is not in an operating mode.

TA ONLY White Indicates that the TCAS is in a traffic advisory (TA)mode only.

RA FAIL Red Indicates that resolution advisories are not available.

TCAS INOP White Indicates that the TCAS is not enabled, but isstrapped.



G. PFD Miscellaneous Annunciators

NOTE: Figure 2-1-23 shows the locations of the PFD miscellaneous annunciators.

(1) Air Data Computer Source Annunciator

The pilot or copilot can select the air data computer (ADC) through the pushbuttonson the reversionary panel. The pushbuttons select either ADC No. 1 or ADC No. 2as the source for air data displays on the PFD. The PFD displays an amber ADCannunciation to indicate which ADC is providing the air data information. Table2-1-36 lists the annunciations and associated colors for the ADC annunciator.

Table 2-1-36. ADC Source Annunciations

SG Reversion Selection ADC ReversionSelection

ADC Annunciation Color

Pilot Copilot Pilot Copilot Pilot Copilot N/A

Normal Normal Normal Normal None None Amber

Normal Normal Reversion Normal ADC2 ADC2 Amber

Normal Normal Normal Reversion ADC1 ADC1 Amber

Normal Normal Reversion Reversion ADC2 ADC1 Amber

Reversion Normal N/A Normal ADC2 ADC2 Amber

Reversion Normal N/A Reversion ADC1 ADC1 Amber

Normal Reversion Normal N/A ADC1 ADC1 Amber

Normal Reversion Reversion N/A ADC2 ADC2 Amber

(2) Air Data Computer Test Annunciator

The PFD displays a red ADC TEST annunciation when calibrated airspeed valid andbarometric altitude valid from the displayed ADC source indicate a functional test.The ADC source annunciation and ADC TEST annunciation share the same locationon the PFD. The ADC TEST annunciation has priority over the ADC sourceannunciation. While the ADC TEST annunciation is displayed, the air datainformation remains displayed on the PFD.

(3) AHRS Test Annunciator

The PFD displays a red ATT TEST annunciation in the upper-center portion of theADI sphere when pitch angle valid and roll angle valid from the displayed headingsource indicate a functional test. When magnetic heading valid from the displayedattitude heading source indicates a functional test, a red HDG TEST annunciation isdisplayed on the full heading compass scale or partial heading compass scale. TheATT TEST and ATT FAIL annunciations share the same locations on the PFD, as dothe HDG TEST and HDG FAIL annunciations. The ATT FAIL and HDG FAILannunciations have priority over the ATT TEST and HDG TEST annunciations.During an AHRS test, the attitude and heading information remain displayed.




260

VOR 1VOR 2ILS 1ILS 2

FMSILS1 HDG

TEST

ATTTEST

ADC TEST

350

235

245

001HDG

359

E

W

21

12

30

15

33

S

24200

260

280

220

.410 M

VOR1

ADF2

YDIASHDG

200 RA

14500

14500

GS

3

1000

1

2

0

1

2

3

25

LOC

29.92 IN

20

20 20

20

10

10 10

10

0020

80240

1

9

AD-51198@

GSAP

143

N3

6

CRS

TGTTTG5MIN

13.1 NM

ADC TESTANNUNCIATOR

AHRS TESTANNUNCIATORS

NAVIGATIONSOURCE

ANNUNCIATOR

Figure 2-1-23 (Sheet 1). PFD Miscellaneous Annunciations




M

1

260

25

260

24240

220

200

280

WDSHEARWSHDG

10 10

20 20

10 10

20 20

SXTK

200 RA.410 M

145 00

14500

14320

8000

29.92 IN

DTK359

HDG001

0

1

23

32

1

CAT2

33N

360

1

9 R2

FMS

14000

3000

1000

VOR1

ADF2

12.5HDR25

TGTGSPD

245KTS

350AOA

ATT2

MAG2

VERTICALTRACK ALERTANNUNCIATOR

WINDSHEARANNUNCIATOR

FMS CROSSTRACKMODE ANNUNCIATOR

AD-51199@

SYMBOLGENERATOR SOURCE

ANNUNCIATOR

SG1SG2

ADC2SG2

AIR DATACOMPUTER SOURCE

ANNUNCIATOR

ADC1ADC2

ILS APPROACHCATEGORY

ANNUNCIATOR

CAT1CAT2

VTA

APPFMS

KDVT

MSG

FMS ACCURACYANNUNCIATOR

FMS STATUSANNUNCIATOR

WPTDR

DGRINTEG

FMS MESSAGEANNUNCIATOR

Figure 2-1-23 (Sheet 2). PFD Miscellaneous Annunciations



(4) Symbol Generator Annunciator

The symbol generator (SG) source annunciation provides an indication to the pilot orcopilot that the display processing function is reverted to a single IC-600 IAC.Symbol reversion forces the PFD formats on the pilot side and copilot side to be thesame. An amber SG1 annunciation is displayed when the SG reversion switch onthe reversionary panel indicates that all SG functions are reverted to the left-sideIAC. An amber SG2 annunciation is displayed when the SG reversion switchindicates that all SG functions are reverted to the right-side IAC. When the SGreversion switch indicates a normal SG selection, no annunciations are displayed oneither PFD.

(5) Navigation Source Annunciator

The PFD displays navigation source annunciations in a left-justified, four-characterfield. The pilot or copilot pushes the NAV button on the DC-550 Display Controller toselect a short range navigation (SRN) source for display on the PFD. The navigationsource annunciation defaults to VOR1 on the pilot’s side or VOR2 on the copilot’sside if the display controller becomes invalid. Table 2-1-37 lists the navigationsource annunciations and associated colors that can be displayed on the PFD. Thenavigation source latches to the last selected source if the display controller isinvalid.

Table 2-1-37. Navigation Source Annunciations

Annunciation Color Primary Navigation Source

VOR1 Green NAV No. 1

VOR2 Green NAV No. 2

ILS1 Green NAV No. 1 (TTL)

ILS2 Green NAV No. 2 (TTL)

FMS Magenta FMS

(6) FMS Cross-Track Mode Annunciator

The PFD only displays the FMS cross-track mode annunciator if the selectednavigation source is an FMS. An amber SXTK annunciation is displayed when theFMS cross-track mode is set. The PFD removes the SXTK annunciation from thedisplay format for invalid FMS modes and FMS warnings.



(7) FMS Accuracy Annunciator

The FMS accuracy annunciations share the same position on the PFD. Theseannunciations are only displayed if the selected navigation source is an FMS. Thepriority for displaying FMS accuracy annunciations is from highest to lowest asfollows:

• APP (GPS Approach Mode)

• No annunciation for the terminal mode

• APP (FMS Approach Mode).

(a) GPS Approach Mode

For a GPS approach, the PFD displays a cyan APP annunciation that flashesfor 10 seconds at a rate of one second on and 0.5 seconds off, then remainssteady. Full scale sensitivity of the cross-track distance/deviation is ±0.3 NM.Full scale sensitivity of the vertical deviation is ±150 feet.

(b) GPS Terminal Area Mode

For a GPS terminal area mode, the PFD displays no annunciation. Full scalesensitivity of the cross-track distance/deviation is ±1 NM. Full scale sensitivityof the vertical deviation is ±500 feet.

(c) FMS Approach Mode

For an FMS approach mode, the PFD displays a magenta APP annunciation.Full scale sensitivity of the cross-track distance/deviation is ±1.25 NM. Fullscale sensitivity of the vertical deviation is ±150 feet.



(8) FMS Status Annunciators

When the selected navigation source is an FMS, the PFD displays FMS statusannunciations for waypoint alert, dead reckoning, degrade, and GPS integrity in thesame location on the display format. Table 2-1-38 lists the FMS statusannunciations in priority order.

Table 2-1-38. FMS Status Annunciations

Annunciation Color Description

WPT Amber The PFD displays this amber annunciationwhen FMS waypoint alert is active.

DR Amber The PFD displays this amber annunciationswhen FMS dead reckoning is active.

DGR Amber The PFD displays this amber annunciationwhen the FMS navigation integrity warningis active.

INTEG Amber The PFD displays this amber annunciationwhen GPS integrity fail is active.

(9) FMS Message Annunciator

The PFD displays an amber MSG annunciation when an FMS message warning isactive. The annunciation flashes continuously when active to alert the pilot orcopilot. The PFD removes the MSG annunciation from the display format for invalidFMS modes and warnings.

(10) Vertical Track Alert Annunciator

The PFD displays an amber vertical track alert (VTA) annunciation above the verticaldeviation scale when the selected navigation source is an FMS. The VTAannunciation is above the FMS label on top of the FMS vertical deviation scale. TheVTA annunciation flashes for five seconds when activated, then remains steady. ThePFD removes the VTA annunciation from the display format for an invalid verticaltrack alert.

(11) ILS Approach Category Annunciator

The PFD displays a steady green CAT1 annunciation if the following criteria issatisfied:

• CAT2 RA minimums is set

• CAT2 is enabled

• The displayed RA minimums value is greater than or equal to 80 feet

• A single PFD is displaying valid ILS deviations

• The remaining CAT2 conditions are not met.



The CAT1 annunciation flashes for five seconds, then remains steady if CAT2 waspreviously displayed before the CAT2 conditions were dropped and reverted to CAT1conditions.

The PFD displays a steady green CAT2 annunciation if all CAT2 conditions are met,and CAT1 was not previously annunciated. The CAT2 annunciation flashes for fiveseconds and then remains steady if CAT1 was previously displayed before the CAT2conditions are met. An amber CAT2 annunciation flashes for 10 seconds and thenremains steady if CAT2 goes invalid after being active. The CAT2 conditions arelisted below:

• No reversions (SG, AHRS, ADC) on either PFD

• Attitude and heading valid on both PFDs

• Valid airspeed and altitude on both PFDs

• No comparison monitors are tripped on either PFD

• An active approach mode

• Glide slope deviation is valid on both PFDs

• CAT2 RA minimums set (greater than 80 feet; less than 200 feet).

(12) Windshear Annunciator

The PFD displays the WDSHEAR annunciation using the color logic specified inTable 2-1-39 if the windshear computer indicates either a windshear caution orwindshear warning condition. For both windshear conditions, the WDSHEARannunciation flashes for 10 seconds at a rate of one second on and 0.5 seconds off,then remains steady. If the WDSHEAR annunciation is currently flashing and theannunciation transitions from a caution status to a warning status, or from a warningstatus to a caution status, then the annunciation flashes for an additional 10 secondsbefore becoming steady.

Table 2-1-39. Windshear Annunciations

Windshear Type Display Color

Caution Amber

Warning Red



H. PFD Comparison Monitor Annunciators

NOTE: Figure 2-1-24 shows the locations of the PFD comparison monitorannunciators.

Comparison monitors indicate to the pilot and copilot that there is a difference betweenpilot and copilot displayed data. All comparison monitors flash for 10 seconds whenactivated, then remain steady.

(1) Indicated Airspeed (IAS) Comparison Monitor Annunciator

The PFD displays an amber IAS annunciation in the upper left corner of airspeedtape when the IAS comparison monitor becomes active. The IAS comparisonmonitor activates when the IAS deviates by more than 10 knots. The IAScomparison monitoring occurs only if both IAS values are above 60 knots.

(2) Pitch (PIT) and Roll (ROL) Attitude Comparison Monitor Annunciator

The PFD displays pitch and roll attitude annunciations within the upper half of theADI sphere. The PFD displays an amber PIT annunciation when the pitchcomparison monitor becomes active, which occurs when the pitch attitude deviatesby more than 5 degrees. The PFD displays an amber ROL annunciation when theroll comparison monitor becomes active, which occurs when the roll attitude deviatesby more than 6 degrees. When both comparison monitors become activesimultaneously, the PFD displays an amber ATT annunciation.

(3) Radio Altitude (RA) Comparison Monitor Annunciator

The PFD displays an amber RA annunciation to the right of the airspeed tape whenthe RA comparison monitor becomes active. The comparison monitor arms when atleast one radio altimeter is less than 2,500 feet. Pushing the system TEST knob onthe DC-550 Display Controller forces the RA annunciation to go off.

(4) Glideslope (GS) and Localizer (LOC) Comparison Monitor Annunciators

The PFD displays the GS, LOC, and ILS annunciations in the same location on thedisplay format, which is just below the radio altitude comparison monitor annunciator.The PFD displays an amber GS annunciation when the glideslope comparisonmonitor becomes active because of the following conditions:

• Glideslope values differ by more than 51 millivolts

• LOC is the selected navigation source.

The PFD displays an amber LOC annunciation when the localizer comparisonmonitor becomes active because of the following conditions:

• Localizer values differ by more than 40 millivolts

• LOC is the selected navigation source.

The PFD displays an amber ILS annunciation when both the glideslope and localizercomparison monitors become active simultaneously.




AP

MSG

M

24

1

260

25

260

240

220

200

280

VSLNAV

10 10

20 20

10 10

20 20

200 RA.410 M

145 00

14500

14320

8000

29.92 IN

DTK359

HDG001

0

1

23

32

133N

360

1

9 R2

FMS

14000

3000

1000VOR1

ADF2

12.5HDR25

TGTGSPD

245KTS

350

AOA RA

MAG2

AD-51200@

FMS

KDVT

HDG

YD

ALT

IAS

LOCCAS

INDICATED AIRSPEEDCOMPARISON MONITOR

ANNUNCIATOR

PITCH/ROLL ATTITUDECOMPARISON MONITOR

ANNUNCIATOR

RADIO ALTITUDECOMPARISON MONITOR

ANNUNCIATOR

GLIDESLOPEAND LOCALIZER

COMPARISON MONITORANNUNCIATOR

CAS COMPARISONMONITOR

ANNUNCIATOR

ALTITUDECOMPARISONMONITORANNUNCIATOR

HEADINGCOMPARISONMONITORANNUNCIATOR

ATT

Figure 2-1-24. Comparison Monitor Annunciators



(5) CAS Comparison Monitor Annunciator

The PFD displays an amber CAS MSG annunciation below the glideslope andlocalizer comparison monitor annunciation area. The CAS MSG annunciation flashesfor the first 10 seconds, then remains steady. The comparison monitor becomesactive when a red level CAS message is output by the on-side IAC, but not thecross-side IAC.

(6) Heading (HDG) Comparison Monitor Annunciator

The PFD displays an amber HDG annunciation below the RA minimums digitalreadout when the heading comparison monitor becomes active. The comparisonmonitor threshold is 10 degrees.

(7) Altitude (ALT) Comparison Monitor Annunciator

The PFD displays an amber ALT annunciation in the upper right corner of thealtitude tape when the altitude comparison monitor becomes active. The altitudecomparison monitor activates when pressure altitudes from the air data computersdeviate by more than 200 feet.

I. PFD Test Mode Operation

The system test provides for a failure mode annunciation and familiarization of the PFDdisplay format. The test is functional while weight-on-wheels is sensed and airspeed isless than 50 knots. Pushing and holding the system TEST knob on the applicable DC-550Display Controller initiates the test mode on the corresponding PFD. The test modecauses the PFD to display a test page format, which remains displayed for the first fourseconds the button is held in. After the first four seconds, an initiated built-in testactivates.

(1) Weight-On-Wheels Test Mode Display Formats

The test page format of the failure mode annunciation and familiarization provide thesymbology and colors illustrated in Figure 2-1-25. Display formats not specificallymentioned do not change. The PFD displays the following items in their invalidstate:

• Altitude select digital readout

• Altitude indicator (tape and digital readout)

• Vertical deviation (scale and pointer)

• Baro set digital readout

• Vertical speed indicator (pointer and digital readout)

• TTG/GSPD/ET readout

• Heading select digital readout

• Lateral deviation (scale and bar)

• Distance readout.




INVALID MACHDIGITAL READOUT(AMBER DASHES)

INVALID AIRSPEEDREFERENCE

DIGITAL READOUT(AMBER DASHES)

INVALID COURSE/DESIRED TRACK

READOUT(AMBER DASHES)

INVALID DISTANCEREADOUT

(AMBER DASHES)

INVALID HEADINGSELECT DIGITAL


RA MINIMUMDIGITAL READOUT

INVALID TTG/GSPD/ET READOUT(AMBER DASHES)

INVALID BARO SETDIGITAL READOUT(AMBER DASHES)

INVALID ALTITUDEINDICATOR(RED X)

INVALID ALTITUDESELECT DIGITALREADOUT(AMBER DASHES)

TEST ANNUNCIATION(MAGENTA)

AP YD

HDGFAIL

I MRAILSCAS

- - -HDG

- - -

E

W

21

12

30

15

33

S

24

.- - -

100

HDG

- - - - -

GS

3

VS

1

2

0

1

2

3

TEST

O

N3

6

CRS

TTG- - -MIN

- - -

ATTFAIL

IAS

ALT

RADIOALTITUDE

VALID

INVALID AIRSPEEDINDICATOR

(RED X)

INVALID VERTICALDIVIATION DISPLAY(RED X)

MARKERBEACONS

HEADING FAILANNUNCIATION(RED)

INVALID VERTICALSPEED INDICATOR(RED BOX ANDDIGITS)

AD-51201@

INVALID LATERALDEVIATON DISPLAY

(RED X)

ATTITUDE FAILANNUNCIATION

(RED)

ATT

WINDSHEAR

AOA

M200 RA

- -.- - IN

- - . - NM

Figure 2-1-25. PFD Familiarization Test Format



• Course select/desired track (digital readout and pointer)

• Mach digital readout

• Airspeed indicator (tape and digital readout)

• Airspeed reference digital readout.

The PFD is forced to display the following items:

• All comparison monitor annunciators

• A red ATT FAIL annunciation in the ADI sphere

• A red HDG FAIL annunciation in the HSI display

• All three beacon markers

• A magenta TEST in the upper-left portion of the display

• An amber AP annunciation in the autopilot status annunciation field

• An amber YD annunciation in the yaw damper status annunciation field

• RA minimum displays at the last set value

• Radio altitude valid display.

The PFD removes the following display items:

• All bugs (airspeed, altitude, heading, drift, vertical speed)

• All pointers (bearing, To/From)

• Flight director information (command bars/cues, mode annunciations, FPAsymbology)

• Low airspeed awareness indicator

• Vspeed bugs and digital readouts

• Vmo/Mmo indicator

• Acceleration trend vector

• Low altitude awareness indicator

• Altitude trend vector

• Vertical speed reference digital readout.

(2) Not Weight-On-Wheels Test Mode Display Formats

The PFD displays a red TEST annunciation in the armed lateral flight directorannunciation field, and removes the following items:

• Flight director modes

• Flight director command bars

• Flight director couple arrow.



J. Multifunction Display Operation

The multifunction display (MFD) provides the pilot with a variety of displays that are menudriven. These menus are controlled by six bezel-mounted menu keys and one rotary knobon the BL-871 MFD Bezel Controller, or by pushbuttons on the DC-550 Display Controller.The menu keys work in concert with the menu selections provided at the bottom of theMFD. Figure 2-1-26 shows the front panel layout for the MFD bezel controller. The menuschange as a function of which mode is selected for the display. The MFD displays include:

• Main menu display

• System page menu display

• MFD menu display

• Electronic checklists

• TCAS data

• Weather radar display

• Weather radar mode annunciations

• Map display for FMS lateral navigation

• Plan display for FMS lateral navigation.

AD-48769@

Figure 2-1-26. BL-871 MFD Bezel Controller



(1) General Operation

The MFD bezel controller menus are organized in the hierarchical structure shown inFigure 2-1-27. The main menu selection is provided on each MFD as the power-updefault menu. A number of submenus can be accessed through the main menudisplay. A white box is placed around the selected menu item. A white box is alsoplaced around momentary menu selections while the corresponding menu key isdepressed. Selecting a new submenu item removes the currently displayed menuand any outline boxes that are being displayed. The newly selected submenu isthen displayed in its previously configured state, or default state if the submenu isbeing selected for the first time.

A C B

AD-51208@

SYS MFD CKLSTM/PRNGTCAS WX

MAPPLAN

MENU

RTN SKP LNBKM/PRNGPAG RCL ENT

CHECK LIST

RTN SKP RCL ENT

JOYSTICK

RTN T/OECSA/I

M/PRNGFUEL HYD ELEC

RTNNAVAPT DATA

M/PRNGSPDS MAINT

MENU

RTNNAVAPT DATA

M/PRNGSPDS JSTK MAINT

MENU

RTNV189 SET

VR---

V2---

AP---

M/PRNG

Figure 2-1-27. MFD Bezel Menu Tree



When the MFD bezel menu indicates MENU INOP, the MFD menu is replaced by anamber MENU INOP annunciation as shown in Figure 2-1-28.

AD-49916@

MENU INOP

Figure 2-1-28. MFD INOP

(a) Menu Operations

Table 2-1-40 describes the menu operations for the following menu operationsthat can be performed through menu keys on the MFD bezel controller:

• Selecting a submenu or subsequent menu page

• Toggling a menu item selection

• Momentarily selecting a menu item

• Selecting a parameter for display

• Selecting a variable parameter for setting a display value.

Table 2-1-40. Menu Key Operations

Operation Description

SubmenuSelections

Pushing the bezel key that is associated with a submenu label selects that submenu,causing the MFD to display the submenu. The RTN menu key on the bezel controllerselects (activates) the main menu.

TogglingSelections

A toggling selection consists of either single or multiple labels associated with a menukey. Pushing the key causes the menu item to sequence through the possibleassociated selections.

MomentarySelections

A momentary action is performed by pushing a menu key associated with a labeldescribing an action that occurs only once per key press, or continues as long as thekey is held down.

ParameterSelections

A parameter or set of parameters is selected for display when the associated menu keyis pushed. This causes the menu item to be boxed and the parameter to be displayedon the MFD in the appropriate format. Repeatedly pushing the menu key does notdeselect the parameter.

Select forSet

The menus provide access to parameters such as the V1 reference. The rotary knob isused for setting the parameter values for display on the MFD.



(b) MFD Rotary Knob

The MFD rotary knob allows parameter values to be entered for display, andcontrols the range of the values. Rotary knob control is accomplished bypushing the menu key associated with the desired parameter. When aparameter is selected, an annunciation is displayed above the rotary knoblabeling its current functionality. The pilot or copilot can then rotate the rotaryknob to adjust the parameter values (select for set). The PFD displays amberdashes instead of the parameter value when an invalid parameter value isselected.

When the map or plan format is selected for the MFD, the pilot or copilot usesthe rotary knob to select various map ranges. Rotating the knob selects amap or plan range in increments of: 2.5, 5, 12.5, 50, 100, 150, 300, and 600nautical miles (NM). The power-up default setting is 100 NM.

(2) MFD Bezel Menus

The MFD bezel menus are described in the following paragraphs. Unless otherwisespecified, all text items are displayed in small white characters and centered abovethe corresponding menu key for a menu selection. The knob label (M/P RNG, SET)also consists of small characters unless otherwise specified. The actual bezel menudisplays do not show the menu headings (i.e., MAIN is not shown above the mainmenu bezel display). However, the dim white line above the menu selections isalways displayed.

(a) MFD Main Menu

The MFD main menu is the power-up default menu, and is displayed uponreturn from any submenu. Figure 2-1-29 shows the main menu bezel display.The M/P RNG menu label is removed from the main menu display if WX datais selected for the MFD. The main menu display is automatically selected ifthe TCAS computer detects a resolution advisory (RA) or traffic advisory (TA)condition. The TCAS display window overrides the system page menu displayon the MFD to display TCAS information for the pilot or copilot.

SYS

AD-51209@

MFD CKLSTM/PRNGTCAS WX

MAPPLAN

MENU

Figure 2-1-29. MFD Main Menu Display



Main menu key selections are described in the following paragraphs:

1 SYS Menu Key

Pushing the SYS menu key selects the system page bezel menu forcontrolling the system pages displayed on the MFD.

2 MFD Menu Key

Pushing the MFD menu key selects the MFD bezel menu for controllingnavigation and speed set functions.

3 CKLST Menu Key

Pushing the CKLST menu key provides the following functions:

• The MFD removes the current system page being displayed

• The MFD removes the TCAS display if it was being displayed

• The MFD displays the electronic checklist

• The MFD displays the checklist bezel menu for controlling the checklistfunctions.

4 TCAS Menu Key

Pushing the TCAS menu key provides the following functions:

• Pushing the menu key toggles the MFD between TCAS ON and TCASOFF; the power-up default is TCAS OFF

• Toggling to TCAS ON displays a box around the TCAS menu item,removes the currently selected system page, and displays the TCASzoom window

• Toggling to TCAS OFF removes the box around the TCAS menu item,removes the TCAS zoom window, and displays the previously selectedsystem page.



5 WX Menu Key

Pushing the WX menu key provides the following functions:

• If the weather radar is on and transmitting, pushing the WX menu keytoggles the MFD between WX ON and WX OFF; the power-up defaultis WX OFF

• If the weather radar is not on, pushing the WX menu key has no effect

• Toggling to WX ON displays a box around the WX menu label, selectsthe MFD map display format if the plan display format was selected,and displays WX data on the MFD

• Toggling to WX OFF removes the box around the WX menu item andremoves the WX data from the MFD.

6 MAP/PLAN Menu Key

Pushing the MAP/PLAN menu key provides the following functions:

• Pushing the MAP/PLAN menu key toggles the MFD between the mapand plan display formats; the power-up default is map

• Toggling to MAP removes the box around the PLAN menu item,displays a box around the MAP menu item, removes the plan format,displays the map format, and selects WX ON if the WX display waspreviously removed from the MFD

• Toggling to PLAN removes the box around the MAP menu item,displays a box around the PLAN menu item, removes the map format,displays the plan format, and selects WX OFF if the WX display iscurrently selected.

7 Rotary (M/P RNG) Knob

The rotary knob controls the range for the MFD map or plan display.When the WX display is selected for the MFD, the rotary knob has nocontrol of weather radar range. The weather radar controller is used toselect the desired weather radar range.



(b) System Page Menu

System pages are displayed full time on the MFD display format, regardless ofwhich system page menu is being displayed. The only time a system page isnot displayed is when the checklist or TCAS menu has been selected fordisplay. Figure 2-1-30 shows the system page menu bezel display.

RTN

AD-51210@

T/OECSA/I

M/PRNGFUEL HYD ELEC

Figure 2-1-30. System Page Menu Display

System page menu key selections are described below:

• RTN (Return) - Pushing the RTN menu key returns the MFD to the mainmenu display.

• T/O (Takeoff) - Pushing the T/O menu key selects the takeoff system pagefor display on the MFD. The T/O menu key has two states: selected andnot selected. If the state is selected, pushing the T/O menu key has noeffect. If the state is not selected, pushing the T/O menu key provides thefollowing actions:

– The state transitions to selected

– The MFD removes the box around the previouslyselected system page menu item

– The MFD displays a box around the T/O menuitem

– The MFD removes the previously selectedsystem page

– The MFD displays the takeoff system page.



• ECS (Environmental Control System) - Pushing the ECS menu key selectsthe environmental control system page for display on the MFD. The ECSmenu key has two states: selected and not selected. If the state isselected, pushing the ECS menu key has no effect. If the state is notselected, pushing the ECS menu key provides the following actions:



– The MFD displays a box the ECS and A/I menuitems


– The MFD displays the environmental controlsystem page.

• FUEL - Pushing the FUEL menu key selects the fuel system page fordisplay on the MFD. The FUEL menu key has three states: selected,reset, and not selected. If the state is not selected, pushing the fuel menukey provides the following actions:



– The MFD displays a box around the FUEL menuitem

– A RESET menu label appears above the FUELlabel


– The MFD displays the fuel system page.

If the state is selected, pushing the fuel menu key provides the followingactions:

– The state transitions to reset

– The MFD removes the box around the FUELmenu item

– The MFD displays a box around the RESETmenu item.



If the state is reset, pushing the fuel menu key provides the followingactions:


– The MFD resets the fuel used quantity readouton the fuel system page to zero

– The MFD removes the box around the FUELmenu item

– The MFD displays a box around the FUEL menuitem.

• HYD (Hydraulic) - Pushing the HYD menu key selects the hydraulic systempage for display on the MFD. The HYD menu key has two states:selected and not selected. If the state is selected, pushing the HYD menukey returns the MFD to the main menu display, and maintains the displayof the hydraulic system page. If the state is not selected, pushing the HYDmenu key provides the following actions:



– The MFD moves the RTN label to a positiondirectly above the HYD menu item

– The MFD displays a box around the HYD menuitem


– The MFD displays the hydraulic system page.

• ELEC (Electrical) - Pushing the ELEC menu key selects the electricalsystem page for display on the MFD. The ELEC menu key has two states:selected and not selected. If the state is selected, pushing the ELECmenu key returns the MFD to the main menu display, and maintains thedisplay of the electrical system page. If the state is not selected, pushingthe ELEC menu key provides the following actions:



– The MFD moves the RTN label to a positiondirectly above the ELEC menu item



– The MFD displays a box around the ELEC menuitem


– The MFD displays the electrical system page.

• Rotary (M/P RNG) Knob - The rotary knob remains functional to control therange for the MFD map or plan display formats. When the WX display isselected for the MFD, the rotary knob has no control of the weather radarrange.

(c) MFD Menu

Figure 2-1-31 shows the MFD menu bezel display with an FMS is installed.Figure 2-1-32 shows the MFD menu bezel display without an FMS installed.The MFD menu key selections are described below:

RTN

AD-51211@

NAVAPT DATA

M/PRNGSPDS JSTK MAINT

MENU

Figure 2-1-31. MFD Menu Display with FMS Installed

RTN

AD-51212@

NAVAPT DATA

M/PRNGSPDS MAINT

MENU

Figure 2-1-32. MFD Menu Display without FMS Installed



• RTN (Return) - Pushing the RTN menu key returns the MFD to the mainmenu display

• SPDS (Speeds) - Pushing the SPDS menu key selects the Vspeedssubmenu display that is used for setting and displaying the Vspeeds

• JSTK (Joystick) - Pushing the JSTK menu key selects the joysticksubmenu display that is used for controlling the designator functions

• APT/NAV (Airport/Navaid) - Pushing the APT/NAV menu key toggles theMFD between displaying airport and navaid symbols on the map or plandisplay format

– The power-up default is not selected

– The toggling sequence is as follows:

• The first toggle sequence is NAV; a single box is shown aroundthe NAV menu label. Navaid symbols and identifiers aredisplayed on the selected display format.

• The second toggle sequence is APT; a single box is shownaround the APT menu label. Only airport symbols and identifiersare displayed on the selected display format.

• The third toggle sequence is APT and NAV; a single box is shownaround both the NAV and APT menu labels. Both airport symbolsand identifiers and navaid symbols and identifiers are displayedon the selected display format.

• The fourth toggle sequence is off. The box is removed from themenu along with all airport symbols, identifiers, navaid symbolsand identifiers from the selected display format.

• DATA - Pushing the DATA menu key toggles the MFD between displayingwaypoint identifiers (IDENTS) on the map or plan display format

– The power-up default is IDENTS ON

– The toggling sequence is IDENTS ON, thenIDENTS OFF

– If the toggle sequence is IDENTS ON, the MFDdisplays a box around the DATA menu item

– If the toggle sequence is IDENTS OFF, the MFDremoves the box around the DATA menu item.The MFD also removes all waypoint identifiersfrom the selected display format.



• MAINT (Maintenance) - Pushing the MAINT menu key provides thefollowing functions on the pilot side only:

– The MFD removes the system page currentlybeing displayed

– If TCAS is selected for display, the MFDremoves the maintenance presentation

– The central maintenance computer (CMC)displays CMC maintenance data on the MFD.

• The MAINT menu key is inhibited if weight on wheels indicates an in airstatus, or if the CMC FAIL message is present on the CAS display.

.• Rotary (M/P RNG) Knob - The rotary knob remains functional to control the

range for the MFD map or plan display formats. When the WX display isselected for the MFD, the rotary knob has no control of the weather radarrange.

(d) Joystick Submenu

The joystick submenu allows the pilot or copilot to control the MFD designatorfunctions. Figure 2-1-33 shows the joystick submenu bezel display. Thejoystick submenu selections are described below:

RTN

AD-51213@

SKP RCL ENT

JOYSTICKM/PRNG

Figure 2-1-33. Joystick Menu Display


• SKP (Skip) - Pushing the SKP menu key provides the following functions:

– On power-up, the designator is co-located withthe present flight plan waypoint position.



– MAP Format

If the designator is co-located with a flight plan waypoint, pushing theSKP menu key skips the designator to the position of the nextwaypoint in the flight plan.

If the designator is skewed off any flight plan waypoint, pushing theSKP menu key skips the tail of the designator line to the nextwaypoint in the flight plan.

– PLAN Format

If the designator is co-located with a flight plan waypoint, pushing theSKP menu key positions the flight plan so the next waypoint isdisplayed over the designator.

If the designator is skewed off any flight plan waypoint, pushing theSKP menu key push skips the tail of the designator line to the nextwaypoint in the flight plan.

• RCL (Recall) - Pushing the RCL menu key provides the following functions:

– On power-up, the designator is co-located withthe present position

– MAP Format

If the designator is co-located with a flight plan waypoint, pushing theRCL menu key positions the designator at the present position of theaircraft, and removes the designator box from the display.

If the designator is skewed off any flight plan waypoint, pushing theRCL menu key positions the designator over the waypoint from whichthe designator line is extended, and removes the designator line fromthe display.

– PLAN Format

If the designator is co-located with a flight plan waypoint, pushing theRCL menu key positions the designator at the present position of theaircraft, and removes the designator box from the display.

If the designator is skewed off any flight plan waypoint, pushing theRCL menu key positions the flight plan so the waypoint to which thetail of the designator line is connected is co-located with thedesignator.

• ENT (Enter) - Pushing the ENT menu key transmits the latitude andlongitude coordinates of the designator to the selected FMS scratchpad asa requested waypoint.



(e) Vspeeds Submenu

The Vspeeds submenu is shown in Figure 2-1-34. The Vspeeds submenuselections are described below:

RTN

AD-51214@

V189 SET

VR---

V2---

AP---

Figure 2-1-34. Vspeeds Menu Display


• V1

The V1 menu key allows the pilot or copilot to set the V1 takeoff referencespeed. On the first selection of the menu key after power-up, threedashes appear under the V1 menu item. Pushing the V1 menu key thefirst time after power-up causes a default digital value of 89 knots toreplace the dashes. Also, two white boxes appear on the menu display.One box is shown around the Vspeed value, and the other box is shownaround the inner box and the V1 menu item.

NOTE: The dual box indicates that the Vspeed is active and selected fordisplay.

The toggling sequence for the V1 menu key is as follows:

– If a dual box is shown around the V1 menu item,pushing the V1 menu key removes the innerbox.

– If a single box is shown around the V1 menuitem, pushing the V1 menu key removes thebox, and also removes the V1 bug and readoutfrom the airspeed tape on the PFD. The state ofthe V1 menu key is latched until a cold start ofthe IC-600 IAC occurs.



• VR

The VR menu key allows the pilot or copilot to set the VR reference takeoffspeed. On power-up, three dashes appear under the VR menu item.Pushing the VR menu key the first time after power-up causes the dashesto be replaced by the greater value of 89 knots, or the value of V1 if V1has been previously set. Also, two white boxes are shown on the menudisplay. One box is shown around the default Vspeed value, and the otherbox is shown around the inner box and VR label.


The toggling sequence for the VR menu key is as follows:

– If a dual box is shown around the VR menuitem, pushing the VR menu key push removesthe inner box.

– If a single box is shown around the VR menuitem, pushing the VR menu key removes the boxand VR bug, and also removes the VR bug andreadout from the airspeed tape on the PFD.The state of the VR menu key is latched until acold start of the IAC occurs.

• V2

The V2 menu key allows the pilot or copilot to set the V2 takeoff referencespeed. On power-up, three dashes appear under the V2 menu item.Pushing the V2 menu key the first time after power-up causes the dashesto be replaced by the greater value of 89 knots, or the value of VR if VRhas been previously set, or the value of V1 if V1 has been previously set.Also, two white boxes are shown on the menu display. One box is shownaround the default Vspeed value, and the other box is shown around theinner box and V2 menu item.


The toggling sequence for the V2 menu item is as follows:

– If a dual box is shown around the V2 menu item,pushing the V2 menu key push removes theinner box.

– If a single box is shown around the V2 menuitem, pushing the V2 menu key push removesthe box, and also removes the V2 bug andreadout from the airspeed tape on the PFD.The state of the V2 menu item is latched until acold start of the IAC occurs.



• AP Pushbutton

The AP menu item allows the pilot or copilot to set the vertical approachreference speed. On power-up, three dashes appear under the AP menuitem. Pushing the AP menu key the first time after power-up causes thedashes to be replaced by the greater value of 89 knots, or the value of V2if V2 has been previously set, the value of VR if VR has been previouslyset, or the value of V1 if V1 was previously set. Also, two white boxes areshown on the menu display. One box is shown around the default Vspeedvalue, and the other box is shown around the inner box and V2 menu item.


The toggling sequence for the AP menu item is as follows:

– If a dual box is shown around the AP menuitem, pushing the AP menu key push removesthe inner box.

– If a single box is shown around the AP menuitem, pushing the AP menu key removes thebox, and also removes the AP bug and readoutfrom the PFD. The state of the AP menu key islatched until a cold start of the IAC occurs.

• Rotary (SET) Knob - The rotary knob allows the pilot or copilot to set theactive Vspeed value in one knot increments if turned slowly, and at afaster rate if turned quickly.

(f) Checklist Menu

The checklist menu is selected from the MFD main menu. The checklist menuallows the user to display customer-defined pages of text on the MFD.However, the checklist menu is only available on one MFD format at a time.Therefore, if the checklist menu is displayed on the pilot’s side, the copilot’schecklist menu is inhibited until the pilot exits his checklist function. Figure2-1-35 shows the checklist menu bezel display.

RTN

AD-51215@

SKP LN BKM/PRNGPAG RCL ENT

CHECK LIST

Figure 2-1-35. Checklist Menu Display



The checklist program holds up to 400 pages of checklist data, with eachchecklist page displaying up to a maximum of 12 lines of data. Six bezelmenu keys are used to control the checklist display. Their functions aredescribed below:


• SKP (Skip) - Pushing the SKP menu key moves the checklist cursor to thenext incomplete procedure/item in the current list. The cursor positionwraps off the end, back to the beginning of the list. If all procedures/itemsin the current list are complete, the skip function has no effect.

• LN BK (Line Back) - Pushing the LN BK menu key moves the checklistcursor to the previous item in the current list. The cursor automaticallywraps to the previous page, or to the last page, if it was at the top of thelist.

• PAG (Page) - When multiple pages exist for the list of procedures/items,pushing the PAG menu key selects the next page. The checklist cursormoves to the first incomplete item on the new page. Pushing the menukey while on the last page, causes a wrap back to the first incomplete itemon the first page of the list. If an incomplete item is not found on the page,then the cursor moves to the first item on the page. If multiple pages arenot present, the PAG menu key has no affect.

• RCL (Recall) - Pushing the RCL menu key moves the cursor to the firstincomplete procedure/item in the current list. If the cursor is already onthe first incomplete procedure/item, the first key push has no effect. If allprocedures/items in the current list are complete, the first key push movesthe cursor to the top of the list.

• ENT (Enter) - Pushing the ENT menu key either selects an index orprocedure, or changes the status of a checklist item from incomplete tocomplete. After the key is pushed, the checklist moves to the nextincomplete procedure/item in the current list.

• Rotary (M/P RNG) Knob - The rotary knob remains functional to control therange for the MFD map or plan display formats. When the WX display isselected for the MFD, the rotary knob has no control of the weather radarrange.



(3) Joystick

The panel mounted joystick controller also provides control of the selected checklistand MFD designator functions.

(a) Checklist Control

When the checklist menu is selected, the pilot or copilot moves the joystick upor down to advance the checklist cursor through the checklist. The pilot orcopilot can also move the joystick left or right to page through the availablechecklists.

(b) Designator Control

The pilot or copilot uses the joystick to control designator movement on theMFD when the checklist function is not selected.

1 Map Mode

In the map mode, the joystick moves the designator symbol as follows:

• Moving the joystick left causes the designator symbol to move left fromits last position

• Moving the joystick right causes the designator symbol to move rightfrom its last position

• Moving the joystick up causes the designator symbol to move up fromits last position

• Moving the joystick down causes the designator symbol to move downfrom its last position.

2 Plan Mode

In the plan mode, the joystick moves the flight plan while the designatorsymbol remains fixed at the center of the plan format as follows:

• Moving the joystick left causes the flight plan to move right from its lastposition

• Moving the joystick right causes the flight plan to move left from its lastposition

• Moving the joystick up causes the flight plan to move down from its lastposition

• Moving the joystick down causes the flight plan to move up from its lastposition.



(4) Common MFD Map/Plan Format Data

Two basic formats are available on the MFD: a partial arc map display and a planformat (North up) display. The primary difference is the base reference or aircrafthome position, and the display of the heading data. In the map display, the aircrafthome position is fixed at the lower center of the 120 degree arc with the headingdata shown in degrees at the top of the display. In the plan format, the homeposition is fixed at the center of a 360 degree display, with true North (N) and an uparrow shown at the top of the display. Electronic checklists, TCAS display, andsystem pages may be displayed in either the MAP or PLAN modes.

This section relates to navigation display information that is common to both theMFD map and plan formats. The positioning of some information is dependent onthe selected display format. The lower portion of the MFD is always reserved fordisplaying the bezel menu and submenu selections. All the format-independentsymbology is shown and labeled in Figure 2-1-36. The functionality of eachdisplayed item is described in the following paragraphs.

(a) Static Air Temperature Display

The static air temperature (SAT) digital readout is based on static airtemperature data from the MADC. The SAT value is displayed in degreesCelsius, rounded to the nearest degree. The readout has a range from -99 °Cto +99 °C with a resolution of one degree. A plus (+) or minus sign (-)precedes values greater than or less than zero.

A white SAT label annunciates the readout. The readout digits are displayedin green. For an invalid static air temperature, three amber dashes (---)replace the readout digits.

(b) Total Air Temperature Display

The total air temperature (TAT) digital readout is based on total airtemperature data from the selected MADC. The TAT value is displayed indegrees Celsius, rounded to the nearest degree. The readout has a rangefrom -99 °C to +99 °C with a resolution of one degree. A plus (+) or minussign (-) precedes values greater than or less than zero.

A white TAT label annunciates the readout. The readout digits are displayedin green. For an invalid total air temperature, three amber dashes (---) replacethe readout digits.

(c) True Airspeed Display

The true airspeed (TAS) digital readout is based on true air speed data fromthe selected MADC. The TAS readout is displayed in knots, rounded to thenearest knot. The digital readout has a range from 0 to 999 knots with aresolution of one knot.

A white TAS label annunciates the readout. The readout digits are displayedin green. For an invalid true airspeed, three amber dashes (---) replace thereadout digits.



NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS.

BEZEL MENU DISPLAY AREA

N

S

33

15

3012

WE

246

21

3

25

+15

300

SATTATTAS

FMSKDVT12.512

NMMIN

360

50 50

31547.0

SYSTEM PAGEDISPLAY AREA

STATIC AIRTEMPERATURE

DISPLAY

DESIGNATORBEARING ANDDISTANCEREADOUT

NAVIGATIONSOURCE

ANNUNCIATOR

WIND SPEEDREADOUT

WX MODEANNUNCIATORS

AD-51230@

CHK EICAS

STABTGTWX-16

+25

WX TILT ANGLEREADOUT

WIND ANGLEARROW

DESIGNATORSYMBOL

TIME-TO-GOREADOUT

WAYPOINTIDENTIFIER

DISTANCE TOWAYPOINTREADOUT

CHECKEICAS

MESSAGE

TOTAL AIRTEMPERATUREDISPLAY

TRUEAIRSPEEDDISPLAY

HALF-RANGERING

RANGEMARKERS

Figure 2-1-36. Common MFD Map/Plan Format Data



(d) Flight Plan Designator Data

Pushing the JSTK menu key on the MFD bezel controller selects the joysticksubmenu which provides the pilot or copilot control of the designator functions.The MFD displays the designator symbol and designator bearing/distancereadouts when the joystick submenu is selected.

1 Designator Symbol

The designator symbol is an unfilled, cyan box with a dashed line thatconnects to the designator’s origination point. Two sources controlmovement of the designator symbology:

• A panel mounted joystick

• The MFD bezel controller via the DC-550 Display Controller.

The designator symbol is set to the present aircraft position (home) as theinitial reference point. Designator movement is then limited to ±2048 NMin both the X and Y directions. In the map mode, the joystick moves thedesignator symbol in an up, down, left, right, or diagonal direction. Anyportion of the designator symbol beyond the outer range ring is masked.In the plan mode, the joystick moves the flight plan while the designatorsymbol remains fixed at the center of the display.

Menu key selections in the joystick submenu display can also position thedesignator symbol as follows:

• Pushing the SKP menu key skips the designator to the next waypoint

• Pushing the RCL menu key recalls the designator to a referencedwaypoint

• Pushing the ENT key transmits the designator’s latitude and longitudecoordinates to the selected FMS scratchpad as a waypoint.

The MFD removes the designator symbol if any of the following conditionsare met:

• The DC-550 Display Controller is invalid

• Heading is invalid

• Present position is invalid.

2 Designator Bearing and Distance Readouts

When the designator symbol is moved from its home position, the MFDdisplays the designator bearing and distance in a digital readout. Thedesignator bearing readout reflects the heading, while the distance readoutreflects the distance from the home position. The designator bearingreadout has a range from 0° to 360° with a resolution of one degree. Thedistance readout has two ranges: 0.0 NM to 99.9 NM with a resolution of0.1 NM, and 100 NM to 2,048 NM with a resolution of one NM. Thereadout digits are displayed in cyan. For an invalid designator, the MFDremoves the bearing and distance readouts from the display.



(e) Range Display

The range display consists of two range markers and a range ring. Rangerings represents the MFD display range selected by the range control. Therange ring boundary is the heading compass arc. The MFD displays ahalf-range ring for the map format, and a full range ring for the plan format.

1 Map Range

The map range display consists of a half-range ring and range markers.The half-range ring is a ±75 degree arc with a radius one-half the radius ofthe partial compass arc. Range markers are displayed at the ends of thehalf-range ring to indicate the selected range. The following ranges aredisplayed: 2.5, 5, 12.5, 25, 50, 100, 150, 300, and 600 nautical miles ifthe flight plan WX mode is not selected. In the flight plan WX mode, thefollowing ranges are displayed: 2.5, 5, 12.5, 25, 50, 150, 250, and 500nautical miles. The power-up default range is 50 NM. The half-range ringand range markers are displayed in white.

For map displays other than weather, the MFD rotary knob controls therange selection via a serial bus from the DC-550 Display Controller. Whenweather radar data is displayed, the WX controller controls the rangeselection via inputs from the MFD rotary knob.

The range selection is invalid if the DC-550 Display Controller is invalid ifweather radar data is not selected for display. If weather radar data isselected for display, then the range selection is invalid if the SCI digitalbus is invalid, or if the weather radar is transmitting invalid data. For aninvalid range selection, the range marker readouts default to 50 NM.

2 Plan Range

The plan range display consists of a full range ring and range markers.The range markers are centered over the left and right edges of the fullrange ring to provide an indication of the selected range. The rangemarkers display the range in nautical miles as follows: 5, 10, 25, 50, 100,200, 300, 600, and 1200 nautical miles. The power-up default range is100 NM. The full range ring and range markers are displayed in white.

The MFD rotary knob controls the range selection via a serial bus from theDC-550 Display Controller. The range selection is invalid if the DC-550Display Controller is invalid. For an invalid range selection, the rangemarkers default to 100 NM.

(f) Weather Radar Data

The MFD format displays four weather radar (WX) status annunciation fieldswhich reflect the settings of the WX receiver/transmitter. A serial controlinterface (SCI) connects the WX receiver/transmitter to the IAC. The SCI buscontains the WX modes and control information to provide a WXreceiver/transmitter status indication on the MFD. The following paragraphsdescribe the annunciations and data related to the weather radar system.



1 WX Mode Annunciator

The WX mode annunciations reflect the on-side WX controller settings.During IAC reversions, the annunciations reflect the settings of the WXcontroller that is on the same side as the operating IAC. Table 2-1-41provides a description of the different mode annunciations and the linenumbers they are displayed in, the top being line No. 1 and the bottombeing line No. 3.

Table 2-1-41. MFD WX Mode Annunciations

DisplayMode Description

Annunciation Color Line No.

FAIL Amber 3 Test mode and faults detected

FPLN Green 3 Flight plan mode

FSBY Green 3 Forced standby (WOW)

GCR Amber 3 Normal WX with ground clutter reduction

GMAP Green 3 Ground map mode

RCT Green 3 Normal WX with REACT

R/T Green 3 WX with REACT and turbulence

STAB Amber 1 Stabilization off

STBY Green 3 Normal standby

TEST Green 3 Test mode and no faults detected

TGT Green 2 Target alert enable

TGT Amber 2 Target alert enabled and Level 3 WX return detected in forward15° of antenna scan

TX Green 3 The WX is transmitting but is not selected for display and is notin the STBY or FSBY mode

TX Amber 3 The WX is transmitting and weight on wheels indicates onground, but WX is not selected for display and is not in the STBYor FSBY mode

VAR Amber 2 Variable gain

WAIT Green 3 Power-up approximately one minute

WX Green 3 Normal WX on and selected for display

WX Amber 3 Invalid WX control bus

WX/T Green 3 Normal WX with turbulence



2 WX Tilt Angle Readout

The WX tilt angle is displayed in a digital readout. The readout range isfrom -16° to 15° with a resolution of one degree. The MFD displays aminus (-) sign in front of values that are less than zero. A degree symbol(°) is displayed after the tilt angle if auto-tilt is not active. If auto-tilt (only aWU-870 receiver/transmitter feature) is active, the character A is displayedin place of the degree symbol.

The MFD removes the tilt angle readout when the WX test mode isselected and faults are detected. The MFD also removes the tilt anglereadout for an invalid tilt angle.

3 WX Fault Codes

The EFIS can display WX receiver/transmitter fault code information. Faultcodes are BCD characters which indicate specific faults/failures within theweather radar system. When the WX receiver/transmitter mode is test andthe WX fault code indicates fail, the MFD removes the WX tilt angle anddisplays the fault codes in amber.

Fault codes can be displayed on the ground or in the air. Weather radarvideo and other annunciations are not affected while operating in thismanner. When more than one fault code is present, the weather radarreceiver/transmitter cycles through the active codes; the EFIS need onlydisplay each fault code as it arrives. Refer to Section 2.5, Weather RadarSystem, for a listing of all WX receiver/transmitter fault codes.

(g) Wind Display

The FMS provides the wind display information. The MFD then displays thewind speed in a digital readout and the wind angle as an arrow. The readoutand arrow are displayed to the left of the heading compass scale. The winddisplay is provided for both SRN and LRN modes.

1 Wind Speed Readout

The MFD displays the wind speed is displayed in knots in a digital readout.The readout has a range from one knot to 255 knots with a resolution ofone knot. The readout digits are displayed in magenta. The MFD removesthe readout for an invalid wind speed, or when the wind speed readoutindicates 0 knots.

2 Wind Angle Arrow

The MFD positions the magenta wind angle arrow to the right of the windspeed readout. The wind angle is given in degrees. An increasing windangle causes a clockwise rotation of the arrow. The arrow rotates aboutits center with 360 degrees of motion. The wind angle arrow is displayedin magenta. The MFD removes the wind angle arrow for an invalid windspeed, or when the wind speed readout indicates 0 knots.



(h) Time-To-Go Readout

The MFD displays the time remaining from the aircraft’s present position to theactive waypoint in a digital readout. The readout has a range from 0 minutesto 512 minutes with a one minute resolution. A white MIN label is displayednext to the readout digits. The readout digits are displayed in magenta. Threeamber dashes (---) replace the digital readout digits if the time-to-go data isinvalid.

(i) Distance To Waypoint Readout

The MFD displays the distance from the aircraft’s present position to the TOwaypoint in a digital readout. The readout has a range from 0.0 NM to 4096NM. The display resolution is 0.1 NM for distances less than 100 NM, andone NM for distances from 100 NM to 4096 NM. A white NM label isdisplayed next to the readout digits. The readout digits are displayed inmagenta. Three amber dashes (---) replace the readout digits if the distancedata is invalid.

(j) Waypoint Identifier

A waypoint identifier designates the next TO waypoint in the flight plan. TheMFD displays the waypoint identifier in a six-character field. The waypointidentifier is displayed in magenta. For an invalid To waypoint, the MFDremoves the waypoint identifier.

(k) Navigation Source Annunciator

The MFD displays the long range navigation (LRN) source that is currentlyproviding the display data. The annunciation is displayed in magenta.

• Single FMS - The MFD displays an FMS annunciation for a single FMSinstallation.

• No FMS - The MFD blanks the navigation source annunciator field whenno FMS is installed.

• Invalid - The MFD removes the navigation source annunciation if no FMSis installed on the aircraft. The navigation source annunciation defaults tothe power-up default annunciation when the DC-550 Display Controller isinvalid.

(l) CHK EICAS Message

The check EICAS message annunciation is displayed when the EICASwrap-around monitor trips. When enabled, the MFD displays an amber CHKEICAS annunciation.



(5) FMS Map/Plan Data

Various map symbols are shown on the map and plan displays. Some symbols areselected through MFD bezel menu key selections. The FMS-supplied mapwaypoints, airports, and navigation aid data (VOR, DME, or co-located VOR/DME)are each represented by a unique symbol. Pilot-defined holding patterns andtop-of-climb (TOC) and top-of-descent (TOD) symbols are also displayed. Figure2-1-37 illustrates the FMS map data symbols along with the appropriate definition.

In the map mode, the FMS map data is masked beyond the outer range arc. Also,the MFD does not display FMS map data beyond 2048 NM from the present position.Map data is referenced to true heading, and is displayed as defined below unless aheading failure occurs or the FMS fails, in which case all map data is removed fromthe MFD. However, the MFD does not remove FMS map data from the display ifheading is invalid in the plan mode.

(a) Waypoint Symbol and Identifier

The waypoint symbol is a four-pointed star positioned at the latitude andlongitude geographic location. This geographic location is referenced to theaircraft’s present position where selected transitions of the flight plan occur.All waypoints are white, except the TO waypoint which is magenta. If thedisplayed range allows, the MFD can display a maximum of 10 waypoints(including the TO waypoint).

White waypoint track lines connect the waypoints in a sequence determined bythe FMS. If a flight plan gap follows a waypoint, then the waypoint track line isnot drawn between that waypoint and the next waypoint in the flight plansequence.

Each waypoint has a 12-character (two six-character lines) identifier. Theidentifier is displayed to the right of a valid waypoint when the togglesequence of the DATA menu key on the MFD bezel controller indicates IDENTON. When a waypoint is a transition point that has no identifier or symbol, thewaypoint is not displayed. Also, any waypoint that is not part of the primaryflight plan is not displayed.

The MFD displays a pseudo-VOR flight plan when the FROM and TO waypointrecord numbers are the same. For a pseudo-VOR flight plan, the inboundcourse line connects to the FROM/TO waypoint, and a fixed length outboundradial is shown as a dashed line in the opposite direction. If subsequentconnected waypoints follow the pseudo-VOR flight plan, then a track lineconnects the pseudo-VOR flight plan to the remaining flight plan.




PLAB1

PLAB2

*PBD01LL01

KDVT

25

APTNAV

SRPKPHX

NAVAIDACKNOWLEDGEANNUNCIATOR

AIRPORTACKNOWLEDGEANNUNCIATOR

WAYPOINTSYMBOL ANDIDENTIFIER

ALTITUDEPROFILE

SYMBOL (TOC,TOD, BOSC)

NAVIGATION AIDSYMBOL AND

IDENTIFIER

AIRPORTSYMBOL AND

IDENTIFIER

- WAYPOINT (MAXIMUM NUMBER: 10)

- NAVAID: DME ONLY MAXIMUM NUMBER IN ANY INDICATION OF VOR/VOR-DME/DME NAVAID SYMBOLS IS 8

- AIRPORT SYMBOL (MAXIMUM NUMBER: 4)

- TOC/TOD (MAXIMUM NUMBER: 2)

TRACK LINE


N

33

306

3

+15

300

SATTATTAS

FMSKDVT12.512

NMMIN

360

50 50

31547.0


AD-51231@

CHK EICAS

STABTGTWX-16

+25

Figure 2-1-37. Map Symbols



(b) Navigation Aid Symbol and Identifier

The navigation aid (NAVAID) symbol (VOR, DME, co-located VOR/DME)symbol is a triangular arrangement of unfilled rectangles which represent theposition of the NAVAID symbol relative to the present position. The MFDdisplays the green NAVAID symbols when the toggle sequence of theAPT/NAV menu key on the MFD menu display indicates NAV is selected. Agreen NAVAID acknowledge annunciation is displayed when NAV is selected.

The MFD can display a maximum of eight NAVAID symbols. Each NAVAIDsymbol has a 12-character (two six-character lines) identifier. The identifier isdisplayed to the right of a valid NAVAID when the toggle sequence of theDATA menu key on the MFD bezel controller indicates IDENT ON.

(c) Airport Symbol and Identifier

The airport symbol is displayed as a cyan circle which represents the positionof the airport relative to the aircraft’s present position. The MFD displays theairport symbols when the toggle sequence of the APT/NAV menu key on theMFD menu display indicates APT is selected. A cyan airport acknowledgeannunciation is displayed when APT is selected.

The MFD can display a maximum of four airport symbols. Each airport symbolhas a 12-character (two six-character lines) identifier. The identifier isdisplayed on to the right of a valid airport symbol when the toggle sequence ofthe DATA menu key on the MFD bezel controller indicates IDENT ON.

(d) Altitude Profile Symbol and Identifier

The altitude profile symbol (TOC/TOD) is displayed as a white diamond whichrepresents the position of the altitude profile relative to the present position.The MFD can display a maximum of two altitude profile symbols. Eachaltitude profile symbol has a 12-character (two six-character lines) identifier.The identifier is displayed to the right of a valid altitude profile symbol whenthe toggle sequence of the DATA menu key on the MFD bezel controllerindicates IDENT ON.



(6) MFD Map Format Display

The MFD map format display is the power-up default display for the MFD. The mapformat display is also selected by pushing the MAP/PLAN menu key on the MFDbezel controller to toggle to the MAP selection. The following paragraphs describethe map format display. Figure 2-1-38 shows the MFD map format display.

(a) Heading Display

The heading display includes a heading compass scale with a stationaryaircraft symbol, a heading readout, and a heading select bug. The AHRSprovides the heading information needed to construct the heading display.The AHRS source is selected on the reversionary panel. The DC-550provides the data inputs for selected heading.

1 Heading Compass Scale

The heading compass scale consists of a white ±60 degree arc. Whitelong tick marks are displayed at 10 degree intervals, and white short tickmarks at intermediate five degree intervals. Digits and cardinalabbreviations are spaced around the inside of the compass arc at30-degree increments. Eight numeric identifiers (3, 6, 12, 15, 21, 24, 30,and 33) are located at 30, 60, 120, 150, 210, 240, 300, and 330 degrees.Four cardinal abbreviations (N, E, S, and W) are shown at 0, 90, 180, and270 degrees. All digits and cardinal abbreviations rotate with the compass.

A white stationary aircraft symbol is displayed at the centerpoint of thecompass arc. The aircraft symbol indicates the aircraft’s position relativeto magnetic north. The compass arc rotates around the aircraft symbol.Increasing values of heading cause a counterclockwise rotation of thecompass arc. For an invalid heading, the MFD displays a red HDG FAILannunciation in the center of the heading display.

2 Heading Readout

The MFD displays a digital readout of the heading when the partialcompass scale is displayed. The heading readout provides an accurateinterpretation of the aircraft heading. A white pointer box positioned at theapex of the compass scale serves as a place holder for the readout. Thebox points to the current heading value.

The readout has a range from 001° to 360° with a resolution of onedegree. Leading zeros are provided for heading values less than 100degrees. The readout digits are displayed in green. For an invalidheading, three amber dashes (---) replace the digits.




PLAB1

PLAB2

*PBD01LL01

KDVT

25

SRPKPHX

AIRCRAFTSYMBOL

HEADINGSELECT

BUG

0.50R

LATERALDEVIATIONREADOUT

WX VIDEODATA

HEADINGREADOUT

MAG2

HEADING SOURCEANNUNCIATOR

DRIFTBUG


MAP MODEHALF-RANGERING

WX VIDEO FAULTANNUNCIATOR

WX

HEADINGCOMPASSSCALEN

33

30

63

+15

300

SATTATTAS

FMSKDVT12.512

NMMIN

360

50 50

31547.0


AD-51232@

CHK EICAS

STABTGTWX-16

+25

Figure 2-1-38. MFD Map Format Display



3 Heading Select Bug

The MFD displays a cyan heading select bug on the compass arc. Theheading select bug is card referenced. The GC-550 Guidance Control Unitcontrols the position of the bug, which has 360 degrees of motion alongthe outside edge of the compass arc. Increasing values of heading selectcause a clockwise movement of the heading select bug.

A cyan off-scale arrow is displayed just above the compass arc if the bugposition is greater than 45 degrees from the current aircraft heading. Thearrow indicates the direction the bug is nearest to the current heading. Foran invalid heading select, the MFD removes the heading select bug andarrow from the display.

(b) Drift Bug

The MFD positions the magenta drift bug on the outer edge of the compassarc. The drift bug moves around the perimeter relative to the angulardifference between FMS computed heading and actual aircraft heading. Thedrift bug is visible within ±60 degrees of the current heading. Increasingvalues of drift cause a clockwise rotation of the bug. The MFD removes thedrift bug for an invalid drift angle or an invalid heading.

(c) Heading Source Annunciator

The MFD displays the heading source annunciations to the right of theheading readout. Heading source annunciations are based on the position ofthe AHRS button on the reversionary panels. Table 2-1-42 lists the headingsource annunciations that are available for display.



Table 2-1-42. Heading Source Annunciations



Pilot PFD Display Copilot PFD Display

Norm Norm Norm Norm DG1(W)

" " HDG1(W)

DG2(W)

" " HDG2(W)

Norm Norm Norm Rev DG1(A)

MAG1(A)

HDG1(A)

DG1(A)

MAG1(A)

HDG1(A)

Norm Norm Rev Norm DG2(A)

MAG2(A)

HDG2(A)

DG2(A)

MAG2(A)

HDG2(A)

Norm Norm Rev Rev DG2(A)

MAG2(A)

HDG2(A)

DG1(A)

MAG1(A)

HDG1(A)

Rev Norm N/A Norm DG2(A)

MAG2(A)

HDG2(A)

DG2(A)

MAG2(A)

HDG2(A)

Rev Norm N/A Rev DG1(A)

MAG1(A)

HDG1(A)

DG1(A)

MAG1(A)

HDG1(A)

Norm Rev Norm N/A DG1(A)

MAG1(A)

HDG1(A)

DG1(A)

MAG1(A)

HDG1(A)

Norm Rev Rev N/A DG2(A)

MAG2(A)

HDG2(A)

DG1(A)

MAG2(A)

HDG2(A)

NOTES:

(A) indicates amber(W) indicates white" " indicates that there is no display when magnetic is displayed and both pilots have selected

their normal heading source.

(d) Weather Radar Display

The pilot or copilot selects the weather radar (WX) display by pushing the WXmenu key on the MFD bezel controller. The weather radar display iscase-referenced. The weather radar (WX) receiver/transmitter providesweather information to the MFD through a video interface bus.

1 WX Video Data

The MFD displays the in a 120° pattern if sector scan has not beenselected on the weather radar controller. If sector scan has been selected,the WX video data is displayed in a 60° pattern. The 60° scan is furtheridentified by two white azimuth marks on the half-range ring at ±30degrees to either side of an imaginary line running through the center ofthe stationary aircraft symbol. Table 2-1-43 specifies the display colors forweather radar and ground map returns.



Table 2-1-43. WX and GMAP Mode Return Colors

Return Weather Mode Groundmap Mode

Level 0 Black Black

Level 1 Green Cyan

Level 2 Yellow Yellow

Level 3 Red Magenta

Level 4 Magenta Black (N/A)

REACT Cyan Black (N/A)

Turbulence White Black (N/A)

A weather radar fault by itself does not remove or prevent weather radarreturns. If the MFD is not receiving weather radar data, it erases theweather video data. The MFD also removes the weather video data for aninvalid video interface bus or an invalid WX range. The MFD clears theweather video data for range changes, transitions into 60° scan, and ontransitions into or out of a system test or the GMAP mode.

2 WX Video Fault Annunciator

Each MFD that is displaying WX video data is monitored for activity toprevent misleading data. A WX video fault occurs when the followingconditions are met:

• The WX mode annunciation on the MFD is WX, GMAP, TEST, or WX/T

• WX is selected for display on the MFD bezel menu

• The WX activity bit is set to valid and set to no activity.

If a WX video fault occurs, the MFD performs the following functions:

• It erases the WX video data from the display

• It displays an amber WX annunciation at the centerpoint of thehalf-range arc, just above the stationary of the aircraft symbol.



(e) Lateral Deviation Readout

The MFD displays the magnitude and direction of the aircraft deviation fromthe desired track in a digital readout above the system page display area. Thereadout has a range from 0.1 NM to 99.9 NM with a resolution of 0.1 NM fordeviations between one and 100 nautical miles. For deviations greater than orequal to 100 nautical miles, the readout has a range from 100 NM to 128 NMwith a resolution of one NM. The readout digits are displayed in white. TheMFD removes the lateral deviation readout for an invalid cross-track deviationor when magnetic heading is invalid.

A white L label is displayed with the deviation readout when the deviation isless than zero, indicating that the aircraft is to the left of the desired track. Awhite R label is displayed with the deviation readout when the deviation isgreater than zero, indicating that the aircraft is to the right of the desired track.The L or R designations are removed when lateral deviation is zero.

(7) MFD Plan Format Display

The aircraft heading is presented on the MFD plan format as a North-up display ofthe active flight plan. The plan format display is selected by pushing the MAP/PLANmenu key on the MFD bezel controller to toggle to the PLAN selection. Headinginformation includes a north-up arrow and an aircraft symbol that rotates with respectto the active flight plan. Figure 2-1-39 shows the MFD plan format display.

In the plan mode, heading data is used to orient the aircraft symbol as it movesaround the active flight plan. A positive increasing angular heading causescounterclockwise rotation of the aircraft symbol. The active flight plan waypoint isdisplayed at the center of the range ring.

(a) North-Up Arrow

The plan mode is always displayed in a true north-up heading format. A whitenorth-up arrow indicates the orientation of magnetic North relative to the planformat. The north-up arrow symbol consists of a segmented arrow with theletter N centered within the arrow. The north-up arrow points to the top of thedisplay format.

(b) Aircraft Symbol

The yellow aircraft symbol represents the aircraft’s actual position relative totrue north and the active flight plan. The MFD positions the aircraft symbol onthe plan format relative to the flight plan waypoints, based on the presentposition from the FMS. The MFD removes the aircraft symbol for any of thefollowing conditions:

• Heading is invalid

• Present position is invalid

• Map data is invalid.




25

+15+25300

SATTATTAS

FMSLLO112.512

NMMIN

TGTTX


1200 1200N

PLAN MODERANGE RING

NORTH UPARROW

PLAN MODEAIRCRAFT SYMBOL

AD-51233@


Figure 2-1-39. MFD Plan Format Display



(8) MFD TCAS Display

In either the map or plan format display, the TCAS display is selected by pushing theTCAS menu key on the MFD bezel controller. When the TCAS controller is set tothe auto mode and the checklist is not activated, the level of traffic also triggers theTCAS display on the MFD. The MFD displays the TCAS auto mode annunciationwhenever the TCAS auto mode is enabled. In the auto mode, the MFD displays awhite TCAS AUTO annunciation.

The MFD TCAS display provides the pilot or copilot with information about thebearing and distance to other Mode-C and Mode-S transponder equipped aircraft inthe area. The MFD displays TCAS data in a zoom window in the lower centerportion of the MFD. The zoom window provides an increased resolution of theintruder traffic in the vicinity of the aircraft, while allowing the pilot or copilot todisplay a map or plan format at a greater range. Figure 2-1-40 shows the TCASdisplay format.

The TCAS display indicates the distance, relative altitude or flight level, verticaldirection of movement for up to 12 of the nearest aircraft. If the TCAS cannotdiscern a bearing from another aircraft due to directional antenna shadowing or afailure, that aircraft is not displayed. However, information concerning that aircraft’sdistance, relative altitude, and vertical direction of movement is provided on theTCAS display if it becomes a threat (TA or RA level) to the TCAS equipped aircraft.

(a) TCAS Mode Annunciator

The MFD displays TCAS mode annunciations based on conditions listed inTable 2-1-44. The TCAS mode annunciations are displayed with the prioritylisted in the table: highest first and lowest last.

Table 2-1-44. PFD TCAS Mode Annunciations

Annunciation Color Condition

TCAS FAIL Amber Indicates that TCAS data is invalid.

TCAS TEST Amber Indicates that the TCAS is undergoing afunctional test.

TCAS OFF White Indicates that the TCAS is not in anoperating mode.

TA ONLY White Indicates that the TCAS is in a trafficadvisory (TA) mode only.

RA FAIL Red Indicates that resolution advisories arenot available.

TCAS INOP White Indicates that the TCAS is not enabled,but is strapped.



AD-51234@

TCAS MODEANNUNCIATOR

TCAS FAILTCAS TESTTCAS OFFTA ONLYRA FAIL

ABSOLUTE ALTITUDEDISPLAY

ABVBLWFL

NO BEARINGTARGET DISPLAY

TCAS RANGE SCALE

PROXMITY ADVISORYSYMBOL

2 NM RANGERING

TA TRAFFICSYMBOL

RA TRAFFICSYMBOL

OFF SCALE TRAFFICSYMBOL


RANGE MARKER

TCAS AUTOMODE

ANNUNCIATOR

N

S

33

15

3012

WE

246

21

3

25

+15+25300

SATTATTAS

FMSKDVT12.512

NMMIN

360

50 50

PLAB1

PLAB2

*PBD01LL01

KDVT

TGTTX-16

STAB TCAS TESTABVFL

MAG2CHK EICAS

RA NO BRGTA NO BRG

-05

-10

12

00

31546.0

6


TCASAUTO

Figure 2-1-40. MFD TCAS Display



(b) TCAS Range Scale

The MFD displays a TCAS range ring as a white ±60 degree arc. A whitestationary aircraft symbol is displayed at the center of the arc. The aircraftsymbol is positioned vertically within the TCAS display window so that whenthe smallest range is selected, three nautical mile range from the center of theaircraft symbol to the bottom of the TCAS window is maintained.

(c) Range Marker

The range marker provides a digital readout of the traffic display range, whichis selected through the RMU in the PRIMUS® II Radio System. The rangemarker is displayed below the end of the right side of the outer TCAS rangering. The readout displays the range in nautical miles from the aircraft positionto the outer range ring. The following ranges can be selected: 6, 10, 25, and50 NM. The readout digits are displayed in white. The readout defaults to sixnautical miles if the TCAS display range is invalid.

(d) 2 NM Range Ring

Whenever the selected range is less than 20 nautical miles, a ring of 12 small,white circles are placed in a radius of 2 nautical miles around the aircraftsymbol. The circles are positioned around the aircraft symbol in 30 degreeintervals, with the circles at the 90, 180, 270, and 360 degree positions havingdiameters twice as large as the other circles. The 360 degree position isoriented above the nose of the aircraft symbol. If the selected TCAS range isgreater than or equal to 20 nautical miles, the MFD replaces the ring with awhite arc that is positioned an equal distance between the aircraft symbol andthe range arc.

(e) TCAS Traffic Symbols

Four types of TCAS traffic symbols, based on threat levels, are displayed asfollows:

• Resolution Advisory (RA) Symbol - The RA symbol is a solid red square.The RA symbol is positioned to indicate the intruder aircraft’s relativebearing and distance from your own aircraft.

• Traffic Advisory (TA) Symbol - The TA symbol is a solid amber circle. TheTA symbol is positioned to indicate a threat aircraft’s relative bearing anddistance from your own aircraft.

• Proximate Advisory (PA) Symbol - The PT symbol is a solid cyan diamond.The PT symbol is positioned to indicate a proximate aircraft’s relativebearing and distance from your own aircraft.

• Other Advisory (OA) Symbol - The OT symbol is a unfilled cyan diamond.The OT symbol is positioned to indicate the other aircraft’s relative bearingand distance from your own aircraft.



(f) Intruder Vertical Speed Indication

Intruder vertical speed is indicated by an arrow positioned to the right of theassociated traffic symbol. The arrow points up for climbing traffic, and downfor descending traffic. The arrow is displayed in the same color as theassociated traffic symbol.

(g) Relative Altitude Display

The relative altitude of an intruder aircraft is displayed as two digits and readin hundreds of feet. The digits are centered above the intruder traffic symbol,preceded by a plus sign (+) if the intruder aircraft is above the aircraft’s ownaltitude. The digits are centered below the intruder symbol, preceded by aminus sign (-) if the intruder aircraft is below the aircraft’s own altitude. Thedigits are displayed in the same color as the associated traffic symbol.

Intruder aircraft with zero relative altitude data are centered below the trafficsymbol without any polarity sign if vertical sensing indicates a descent.Intruder aircraft with zero relative altitude data are centered above the trafficsymbol without any polarity sign if vertical sensing indicates a climb. If therelative altitude exceeds ±9900 feet, then a ?? annunciation is displayedinstead of the altitude.

(h) Absolute Altitude Display

Absolute altitude is displayed in the relative altitude position, but uses a threedigit flight level format (example: 23,500 feet is displayed as 235). Absolutealtitude is displayed instead of relative altitude of the traffic aircraft if thefollowing conditions are met:

• No RAs or TAs are displayed

• The TCAS display command is valid

• Aircraft barometric altitude is valid.

Positive values of absolute altitude have no polarity sign. Negative values aredisplayed with a minus sign (-) (i.e., -2100 feet = 21-). The digits aredisplayed in the same color as the associated traffic symbol.

When absolute altitude is selected, FL is displayed in white below the whiteABV/BLW annunciation. If the TCAS intruder altitude reaches ±12,700 feet,then "??" is displayed instead of the altitude. Absolute altitude is onlydisplayed for PA and OA traffic symbols. Relative altitude is displayed for TAand RA traffic symbols.



(i) Off Scale Traffic Symbol

Those RA or TA traffic symbols that have moved beyond the displayed rangeor behind a display stroke mask are shown as half symbols. Half symbols areplaced at the edge of the TCAS writing area and at the correct bearing relativeto the aircraft’s own position. Off-scale traffic symbols have a data tagshowing the relative altitude and a vertical speed sense arrow. The PA andOA traffic symbols are not displayed when out of range.

(j) No Bearing Intruder Display

The MFD presents the no bearing target display if an RA and TA target isencountered that does not have a bearing available for display. The first lineof the display contains a red RA NO BRG message for an RA target withoutbearing information. The second line contains an amber TA NO BRGmessage for a TA target without bearing information.

(9) MFD Checklist Display

In either the map or plan format display, the checklist display is selected by pushingthe CKLST menu key on the MFD bezel controller. The checklist function allows upto 400 customer-defined pages of text to be displayed on the MFD. Independentchecklists are available on both the pilot’s and copilot’s MFD. The checklist page isdisplayed in a window in the lower center portion of the MFD. Figure 2-1-41 showsthe checklist display format.

The checklist page accommodates a 312 character block (12 lines of 26 mediumASCII characters). All text is stroke written for sunlight readability. The checklistoverwrites all information previously displayed in the checklist area and the systempage area. The panel mounted joystick can be used to manipulate the checklist.The checklist cursor is drawn as a box around the selected checklist item. Thedefault color of completed checklist procedures/items are modified to distinguishthem from the incomplete procedures/items. Checklist procedures/items areassigned the colors listed in Table 2-1-45.



CURSOR BOX

AD-51235@

CHECKLISTDISPLAYWINDOW

CHECKLISTTITLE LINE

25

FMSLLO112.512

NMMIN

TGTTX

50 50N

1MASTER INDEX

2

3

EMERGENCYPROCEDURESNORMALPROCEDURESABNORMALPROCEDURES

1/1

23040.0

+15+25300

SATTATTAS


Figure 2-1-41. MFD Checklist Display



Table 2-1-45. Checklist Color Assignments

Checklist Item Color

Cursor Box White

Default Incomplete Item or Procedure Cyan

Complete Item or Procedure Green

"Checklist Unavailable" Message Amber

"Emergency Procedure Complete" Message Magenta

"Abnormal Procedure Complete" Message Magenta

All Disclaimer Text Green

Default Header Text/Page Numbers White

"FROM" Waypoint (Line No. 2) Amber

"TO" Waypoint (Line No. 3) Magenta

Additional Waypoints (Line No. 4 - Line No. 11) Cyan

Present Position (Line No. 12) Green

(a) MFD Bezel Menu Controls

The checklist menu keys on the MFD bezel controller provide the controlinputs for moving the checklist cursor and selecting various procedures anditems on the checklist pages. These menu keys function as follows:

1 Return Function

Pushing the RTN menu key deactivates the checklist display and returnsthe MFD to the main bezel menu.

2 Skip Function

Pushing the SKP menu key moves the cursor to the next incompleteprocedure/item in the current checklist. When the cursor reaches the endof the list, it wraps back to the beginning. If all the procedures/items in thecurrent list are complete, the skip function has no effect.

3 Line Back Function

Pushing the LN BK menu key moves the checklist cursor to the previousitem in the current list. The cursor automatically wraps to the previouspage, or to the last page, if it was at the top of the list.



4 Page Function

When multiple pages exist for the currently displayed list ofprocedures/items, pushing the PAG menu key displays the next page. Thecursor moves to the first incomplete procedure/item on the new page.While on the last page, pushing the PAG menu key causes the cursor towrap back to the first incomplete procedure/item on the first page. If anincomplete procedure/item is not found on the page, then the cursor movesto the first procedure/item on the page. This function has no effect ifmultiple pages are not present.

5 Recall Function

Pushing the RCL menu key moves the checklist cursor to the firstincomplete procedure/item in the current list. If the cursor is already onthe first incomplete procedure/item, the first push of the RCL menu key hasno effect. If all procedures/items on the current list are complete, the firstpush of the RCL menu key moves the cursor to the top of the list.

6 Enter Function

Pushing the ENT menu key either selects an index or procedure, orchanges a checklist item’s status from incomplete to complete. The cursormoves to the next incomplete procedure/item in the current list.

(b) Joystick Controls

When the checklist menu is selected, the pilot or copilot moves the joystick upor down to advance the checklist cursor through the checklist. The pilot orcopilot can also move the joystick left or right to page through the availablechecklists.

1 Joystick Up Function

Moving the joystick up moves the cursor to the previous procedure/item inthe current list. The cursor automatically wraps to the previous page, orthe last page, if the cursor was at the top of the list.

2 Joystick Down Function

Moving the joystick down moves the cursor to the next procedure/item inthe current list. The cursor automatically wraps to the next page, or thefirst page, if the cursor was at the bottom of the list.

3 Joystick Right Function

Moving the joystick right displays the previous checklist/index page. Thecursor moves to the first procedure/item on the newly displayed page. Ifthe first checklist/index page is being displayed, then the lastchecklist/index page is displayed when the function is selected.



4 Joystick Left Function

Moving the joystick left displays the next checklist/index page. The cursormoves to the first procedure/item on the newly displayed page. If the lastchecklist/index page is being displayed, then the first checklist/index pageis displayed when the function is selected.

(c) Checklist Modes

The checklist function has nine modes with each index, checklist, or listingconsidered a mode. These modes are then grouped as a major or minormode as follows:

• Major modes: master index, normal procedure index, normal checklist,emergency procedure index, emergency checklist, abnormal proceduresindex, and abnormal checklist

• Minor modes: waypoint listing, preamble, and disclaimer listing.

An index mode displays the procedures that a pilot or copilot may wish toreview during flight. A checklist mode displays the details of a particularprocedure that a pilot or copilot may wish to check during flight operations.

1 Disclaimer Mode

The disclaimer mode is entered when the checklist function is initializedafter power-up. This mode consists of text explaining to the user thatusing the electronic checklist does not relieve the user from complying withthe checklist contained in the Aircraft Flight Manual. Figure 2-1-42 showsan example of a disclaimer page.

After cycling through all disclaimer pages, the active mode changes to themaster index if it exists, otherwise the normal index mode is entered.When the disclaimer mode is completed, the disclaimer selection in thechecklist is changed to a complete status.

Each time the disclaimer mode is activated, the status of all normalprocedures and checklist items and all abnormal and emergency checklistitems are set to incomplete. While in the disclaimer mode, the pagefunction is the only control function available to progress through thedisclaimer pages. If the disclaimer mode is entered a second time, it canonly be exited by pushing the recall menu key twice.

2 Master Index Mode

The master index mode is a customer option. The master index modecontains listings for the normal procedures index, emergency proceduresindex, and abnormal procedure index. If the master index mode isavailable, it is entered by pushing the CKLST menu key on the MFD bezelcontroller. All selections of the CKLST menu key, excluding the first timeafter a power-up, activate the master index mode. All selections in themaster index are displayed as incomplete.



DISCLAIMER 1/1

USE OF THIS CHECKLIST DOES

NOT RELIEVE THE FLIGHT

CREW OF ITS RESPONSIBILITY

TO COMPLY WITH THE

AIRWORTHINESS AUTHORITY

APPROVED CHECKLIST

CONTAINED IN THE AIRCRAFT

FLIGHT MANUAL/SUPPLEMENTS

AD-51236@

Figure 2-1-42. Disclaimer Page

3 Normal Procedures Index Mode

The normal procedures index contains a list of procedure which can beselected to access checklists for routine flight operations. When achecklist is completed, the corresponding procedure in the normal index isalso denoted as being complete. Figure 2-1-43 shows an example of anormal procedures index page.

The normal index mode is entered from the master index mode if it exists.Otherwise the normal index mode is entered from the disclaimer mode.When the normal index mode is entered, the checklist cursor moves to thefirst incomplete procedure in the list. Any selection from the normal indexmode activates the normal checklist mode for that selection, except thedisclaimer or waypoint listing functions. When the checklist items for aselected procedure are complete, the procedure’s completion status ischanged to complete. The status for a normal procedure and allassociated checklist items changes to incomplete upon re-entering thechecklist mode for that procedure.



1/2

AD-51237@

NORMAL

PROCEDURES

WAYPOINT

LISTING

DISCLAIMER

BEFORE START

ENGINE START

BEFORE TAXI

TAXI

BEFORE TAKEOFF

T/O SPD FLPS 20

T/O SPD FLPS 7

1

2

3

4

5

6

7

8

9

Figure 2-1-43. Normal Procedures Index Page

4 Waypoint Listing Mode

The waypoint listing mode displays the aircraft’s current position, theFROM waypoint, the TO waypoint, and any additional FMS or LRNwaypoints on the flight plan up to a maximum of eight. Line 1 of thedisplay contains the waypoint listing header consisting of the ASCII columnlabels "ID LAT ON XXX". The XXX represents the current navigationsource. Dash marks replace the respective invalid digits if the waypointinformation is invalid or unavailable. Table 2-1-46 lists the data that isdisplayed for each waypoint. Figure 2-1-44 shows an example of awaypoint listing page.

Table 2-1-46. Waypoint Listing Data

Waypoint Data Display Attributes

Identifier Four Characters

Latitude Hemisphere Indication N - Northernor

S - Southern

Digital Latitude XX°XX.X

Longitude Hemisphere Indication W - Westernor

E - Eastern

Digital Longitude XXX°XX.X



AD-51238@

N 33 26.7

N 34 39.1

KPHX

KPRC

KFLG

TO+2

TO+3

TO+4

TO+5

TO+6

TO+7

TO+8

N 35 08.3

W112 00.5 W112 25.2 W111 40.7

PPOS N 32 04.4 W110 45.7

____ __._ ____ __._

____ __._____ __._

____ __._

____ __._

____ __._

____ __._

____ __._

____ __._

____ __._

____ __._

____ __._

____ __._

ID LAT LON

Figure 2-1-44. Waypoint Listing Page

5 Preamble Mode

The preamble mode, if defined by the customer, is a page showing thecharacteristics of the checklist. While in the preamble mode, the onlyactive selections are the page and recall functions.

6 Normal Checklist Mode

The normal checklist mode is entered when any procedure is selected fromthe normal procedures index, except the disclaimer or waypoint listingfunctions. A normal checklist contains specific actions for normal flightoperations. Figure 2-1-45 shows an example of a normal checklist page.

Each procedure listed in the normal procedures index has its own list ofchecklist items which are displayed during the normal checklist mode.When the normal checklist mode is entered, the checklist cursor moves tothe first incomplete checklist item in the list. As each checklist item iscompleted, pushing the enter menu key changes the item’s default color toindicate a complete status. When the last checklist item has beencompleted for that procedure, the normal procedures index mode isactivated and the checklist cursor moves to the next incomplete procedurein the index.



AD-51239@

WHEEL

OVERHEAT

1/1

-190 KIAS

-DOWN

SPEED BELOW

LANDING GEAR

MAINTAIN GEAR DOWN TILL

LIGHT IS OUT AND FOR NO

LESS THAN 10 MIN

CAUTION :

TIRES MAY BE DEFLATED

MAKE A SOFT LANDING

Figure 2-1-45. Normal Checklist Page

7 Emergency/Abnormal Procedures Index Modes

Emergency and abnormal procedures index modes are the same as thenormal procedures index, except they apply to procedures that areemergencies or atypical for a normal operation. The emergency andabnormal index modes contain lists of their respective procedures.

When the emergency or abnormal index mode is entered, the checklistcursor moves to the first procedure in the list. The emergency checklistmode is active whenever a procedure is selected during the emergencyindex mode. The abnormal checklist mode is active whenever a procedureis selected during the abnormal index mode. Procedures displayed duringthe emergency and abnormal index modes always have a status ofincomplete. Figure 2-1-46 shows an example of an emergency proceduresindex page.



1/3

AD-51240@

1

2

3

4

5

6

EMERGENCY

PROCEDURES

ENGINE FIRE

FIRE APU

BAG COMP

SMOKE/FIRE

WHEEL

OVERHEAT

ELECTRIC

SMOKE/FIRE

AIR COND'G

SMOKE

Figure 2-1-46. Emergency Procedures Index Page

8 Emergency/Abnormal Checklist Modes

Emergency and abnormal checklist modes are the same as the normalchecklist mode, except they apply to checklists that are emergencies oratypical for a normal flight. Each of the procedures listed in theemergency or abnormal index mode has its own list of checklist items thatare displayed during the emergency or abnormal checklist mode. Whenthe emergency or abnormal checklist mode is entered, the checklist cursormoves to the first incomplete item in the list. As each checklist item iscompleted, pushing the enter menu key changes the item’s default color toindicate a complete status. The completion status for each of the checklistitems is maintained until the disclaimer mode is activated, or until all of theitems in the list have been completed.

When the last checklist item has been completed for an emergencyprocedure, the display is cleared, the message "EMERGENCYPROCEDURE COMPLETE" is displayed, and the statuses of all checklistitems in that procedure are set to incomplete. Selecting the enter functionwhile the message "EMERGENCY PROCEDURE COMPLETE" is beingdisplayed, clears the message and activates the master index if it exists;otherwise the normal procedures index mode is entered.

When the last checklist item has been completed for an abnormalprocedure, the display is cleared, the message "ABNORMAL PROCEDURECOMPLETE" is displayed, and the status of all the checklist items in thatprocedure are set to incomplete. Selecting the enter function while themessage "ABNORMAL PROCEDURE COMPLETE" is being displayed,clears the message and activates the master index if it exists; otherwisethe normal procedures index mode is entered.



(10) System Page Displays

System pages are displayed full-time on the MFD display format, regardless of whichsystem page is selected through system page bezel menu. The only time a systempage is not displayed is when the checklist or TCAS menu has been selected fordisplay.

The EFIS uses data from the selected channel in each DAU to display engineinformation. If an exceedance occurs while the system page is being displayed, itcauses the box around the digital readout to flash for five seconds at a rate of onesecond on and 0.5 seconds off, then remain steady. If a parameter is currently inexceedance and it enters a higher level exceedance, the box flashes for five secondsand then remains steady. The box changes to the exceedance color but does notflash if a parameter currently in exceedance enters a lower level exceedance.

If a parameter is currently in exceedance with its box flashing and it enters a higherlevel exceedance, the box changes to the new exceedance color and flashes for theremainder of the lower level exceedance’s five second duration. If a parameter iscurrently in exceedance with its box flashing and it enters a lower level exceedance,the box changes to the new exceedance color and flashes for the remainder of thehigher level exceedance’s five second duration. The box is immediately removed if aparameter becomes invalid while in an exceedance with the box flashing. If aparameter is invalid and then becomes valid within an exceedance range, the boxand digits change to the exceedance color and the box flashes for five seconds andthen remains steady.

(a) Electrical System Page Display

The electrical system page format contains display information that provides aschematic representation of the electrical system for the pilot or copilot. Theelectrical system page format displays the dc buses, bus connections, voltageand current readouts for the generators, auxiliary power unit, ground powerunit, and the battery status. Figure 2-1-47 shows the display format for theelectrical system page.

1 Generator Data

The generator data consists of voltage and current readouts for generators1 thru 4. Voltages and currents are displayed inside a schematicrepresentation of the generators. White GEN1, GEN2, GEN3, and GEN4labels annunciate the generator data displays, respectively.

a Generator Voltage Readout

The generator voltage is displayed in a digital readout outlined by awhite box. The readout has a range of 0.0 volts to 40.0 volts with aresolution of 0.1 volts. The readout digits are displayed in green fornormal conditions. A white V label is displayed next to the readoutdigits. For an invalid generator voltage, three amber dashes (--.-)replace the readout digits.



C

00.0GEN4

28.500.000.028.528.5000120000120120

GEN2GPUAPUGEN3GEN1

28.528.57040 C

VA

VA

VA

V

V VESS1 ESS2BATT1 BATT2

GENERATORNO. 1 AND NO. 3

VOLTAGE ANDCURRENT READOUT

AUXILLARY POWERUNIT VOLTAGEAND CURRENT

READOUT

GROUND POWERUNIT VOLTAGEAND CURRENT

READOUT

GENERATORNO. 2 AND NO. 4VOLTAGE ANDCURRENT READOUT

RIGHT DCBUS LINE

LEFT DCBUS LINE

LEFTESSENTIAL

DC BUS LINE

RIGHTESSENTIALDC BUS LINE

BATTERYNO. 1 AND NO. 2VOLTAGE ANDTEMPERATURE

READOUT

CENTER DCBUS LINE

LABELS/UNITS

NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS. AD-51578@

Figure 2-1-47. Electrical System Page Format - Normal Conditions

A generator voltage exceedance causes the outline box and readoutdigits to change colors. Generator voltage exceedances are basedsolely on the existence of a CAS message, as opposed to being withina specified range. If the associated CAS message is generated, thereadout digits and outline box are displayed in amber.

b Generator Current Readout

The generator current is displayed in a digital readout outlined by awhite box. The readout has a range of 0 amps to 723 amps with aresolution of five amps. The readout digits are displayed in green fornormal conditions. A white A label is displayed next to the readoutdigits. For an invalid generator current, three amber dashes (---)replace the readout digits.

A generator current exceedance causes the outline box and readoutdigits to change colors. Generator current exceedances are basedsolely on the existence of a CAS message, as opposed to being withina specified range. If the associated CAS message is generated, thereadout digits and outline box are displayed in amber.



2 Auxiliary Power Unit Data

Auxiliary power unit (APU) data consists of voltages and current readouts.Voltages and currents are displayed inside a schematic representation ofthe APU. A white APU label annunciates the APU data display.

The APU labels and schematic representation are removed from thedisplay when the APU master switch indicates off if the APU turbine speedis invalid. The APU labels and schematic representation are also removedfrom the display if the APU master switch is invalid and the APU turbinespeed is less than 10%.

a APU Voltage Readout

The APU voltage is displayed in a digital readout outlined by a whitebox. The readout has a range of 0.0 volts to 40.0 volts with aresolution of 0.1 volts. The readout digits are displayed in green fornormal conditions. A white V label is displayed next to the readoutdigits. For an invalid APU voltage, three amber dashes (--.-) replacethe readout digits.

An APU voltage exceedance causes the outline box and readout digitsto change colors. The APU voltage exceedances are based solely onthe existence of a CAS message, as opposed to being within aspecified range. If the associated CAS message is generated, thereadout digits and outline box are displayed in amber.

b APU Current Readout

The APU current is displayed in a digital readout outlined by a whitebox. The readout has a range of 0 amps to 723 amps with aresolution of five amps. The readout digits are displayed in green fornormal conditions. A white A label is displayed next to the readoutdigits. For an invalid APU current, three amber dashes (---) replacethe readout digits.

An APU current exceedance causes the outline box and readout digitsto change colors. The APU current exceedances are based solely onthe existence of a CAS message, as opposed to being within aspecified range. If the associated CAS message is generated, thereadout digits and outline box are displayed in amber.



3 Ground Power Unit Data

Ground power unit (GPU) data consists of voltage readouts. Voltages aredisplayed inside a schematic representation of the GPU. A white GPUlabel annunciates the GPU data display.

The GPU voltage is displayed in a digital readout outlined by a white box.The readout has a range of 0.0 volts to 40.0 volts with a resolution of 0.1volts. Readout digits are displayed in green for normal conditions. A whiteV label is displayed next to the readout digits. If weight on wheelsindicates an in air status, then the GPU labels and digital readout areremoved from the display. For an invalid GPU voltage, three amberdashes (--.-) replace the readout digits. No exceedance state exists forGPU data.

4 Battery Data

The DAU provides the battery voltage and temperature data. This data isdisplayed in digital readouts. White BATT1 and BATT2 labels annunciatethe battery data displays.

a Battery Voltage Readout

The battery voltage is displayed in a digital readout outlined by a whitebox. The readout has a range of 0.0 volts to 40.0 volts with aresolution of 0.1 volts. Readout digits are displayed in green fornormal conditions. A white V label is displayed next to the readoutdigits. For an invalid generator voltage, three amber dashes (--.-)replace the readout digits.

A battery voltage exceedance causes the outline box and readoutdigits to change colors. Battery voltage exceedances are based solelyon the existence of a CAS message, as opposed to being within aspecified range. If the associated CAS message is generated, thereadout digits and outline box are displayed in amber.

b Battery Temperature Readout

The battery temperature is displayed in a digital readout outlined by awhite box. The readout has a range of -40 °C to +150 °C with aresolution of one degree. Readout digits are displayed in green fornormal conditions. A white °C label is displayed next to the readoutdigits. For an invalid generator voltage, three amber dashes (---)replace the readout digits.

A battery temperature exceedance causes the outline box and readoutdigits to change colors. Battery temperature exceedances are basedsolely on the existence of a CAS message, as opposed to being withina specified range. If the associated CAS message is generated, thereadout digits and outline box are displayed in red.



5 DC Bus Lines

Five schematic representations are displayed for the left, center, right, leftessential, and right essential dc buses. The DAU sends the current statusof each bus to the IAC for display on the electrical system page. Fornormal conditions, dc bus lines are displayed in green. If an associatedCAS message is generated, the left and right dc bus lines and the left andright essential dc bus lines change to amber to annunciate the condition.For an invalid dc bus, the respective dc bus line changes to amber. If theDAU bus becomes invalid, amber dashes are displayed instead of the buslines.

6 Test Mode Display

The test mode provides the user with a failure mode annunciation andfamiliarization of the electrical system page. The test mode is functionalwhile weight-on-wheels and the airspeed is less than 50 knots. Pushingand holding the TEST button on the appropriate DC-550 Display Controllerinitiates the test mode if the electrical system page was being displayed.The MFD then displays the test page format as long as the test mode isactive. Figure 2-1-48 shows the test mode format for the electrical systempage.

C

- - . -GEN4

- - . -- - . -- - . -- - . -- - . -- - -- - -- - -- - -- - -

GEN2GPUAPUGEN3GEN1

- - . -- - . -- - -- - - C

VA

VA

VA

V

V VESS1 ESS2BATT1 BATT2

GENERATORNO. 1 AND NO. 3

VOLTAGE ANDCURRENT READOUT

(AMBER DASHES)

AUXILLARY POWERUNIT VOLTAGEAND CURRENT


GROUND POWERUNIT VOLTAGEAND CURRENT


GENERATORNO. 2 AND NO. 4VOLTAGE ANDCURRENT READOUT(AMBER DASHES)

DC BUS LINES(AMBER)

BATTERYNO. 1 AND NO. 2VOLTAGE ANDTEMPERATURE



DC BUS LINES(AMBER)

Figure 2-1-48. Electrical System Page - Test Mode



(b) Hydraulic System Page Display

The hydraulic system page format contains display information in the form ofscales and digital readouts which provide identification and separation of thevarious hydraulic parameters. Headers and header lines are displayed inwhite. Figure 2-1-49 shows the display format for the hydraulic system page.

HYDRAULICS

QTY

PRESS

BRAKES

TEMP

IBOB IB OBELEC PUMP

SYS 2

3000

OFF

SYS 1

3000

ON

HEADERS ANDHEADER LINES

HYDRAULICFLUID QUANTITY

SCALES

BRAKETEMPERATUREVERTICALSCALES

HYDRAULICPRESSUREREADOUTS

HYDRAULICELECTRICAL

PUMP ANNUNCIATORS


Figure 2-1-49. Hydraulic System Page Format

1 Hydraulic Fluid Quantity

The DAU provides the left and right hydraulic fluid quantities to the IAC.These quantities are then displayed as an analog scale for each hydraulicsystem. Each scale has a range of 0 to 6.3 quarts. A white QTY label isdisplayed between the two scales.

Each scale consists of two bands and a pointer. The amber left bandrepresents quantities from 0 to 1.1 quarts. The green right bandrepresents quantities from 1.1 to 6.3 quarts. The pointer moves linearlyalong the bottom of the scale between the left and right bands. Thepointer reflects the color of the band to which it is pointing. When ahydraulic fluid quantity value exceeds the limits of a particular band, thepointer changes color. When a hydraulic fluid quantity value exceeds thelimits of the scale, the pointer parks itself at the respective end of thescale. For an invalid hydraulic fluid quantity, the MFD removes the pointer.



2 Hydraulic Pressure

The DAU provides hydraulic pressure data to the IAC. The hydraulicpressure for each hydraulic system is then displayed in a digital readout.Each readout has a range of 0 psi to 5200 psi with a resolution of 100 psi.A white PRESS label is displayed between the two readouts.

The digits are displayed in green for normal pressure conditions between1200 psi and 3400 psi. When a hydraulic pressure exceeds 1200 psi or3400 psi, an amber exceedance box is drawn around the readout and thedigits are displayed in amber. If weight on wheels indicates an on groundstatus and the on-side N2 value is less than 50%, the hydraulic pressure isdisplayed as normal regardless of its current value. For an invalidhydraulic pressure, four amber dashes (---) replace the readout digits.

3 Hydraulic Electrical Pump Status

The status of each hydraulic electrical pump is displayed in a digitalreadout to indicate whether the hydraulic engine is on or off. A green ONannunciation is displayed if the hydraulic electrical pump switch indicatesnormal pressure. A green OFF annunciation is displayed if the hydraulicelectrical pump switch indicates low pressure. A white ELEC PUMP labelis displayed between the digital readouts. For an invalid hydraulicelectrical pump status, three amber dashes (---) replace the digits in thecorresponding readout.

4 Brake Temperature

Brake temperatures are displayed as two vertical scales. The scalesprovide the temperature indication for the outboard and inboard brakes oneach side of the aircraft. Each scale has a range of 0 °C to 1000 °C. Thebottom of the scale represents 0 °C, and the top represents 1000 °C.White OB and IB labels are displayed below each scale to indicate theoutboard and inboard brakes.

The scales consist of two vertical bands and four pointers. The greenlower band represents temperatures from 0 °C to 200 °C. The amberupper band represents temperatures from 200 °C to 1000 °C. The leftoutboard brake temperature pointer is positioned on the left edge of the leftvertical scale. The left inboard brake temperature pointer is positioned onthe right edge of the left vertical scale. The right outboard braketemperature is positioned on the left edge of the right vertical scale. Theright inboard brake temperature pointer is positioned on the right edge ofthe right vertical scale.

The pointers move linearly between the top and bottom of the scales toindicate the brake temperature. The pointers reflect the color of the bandto which they are pointing. When any brake temperature exceeds thelimits of a particular band, the corresponding pointer changes color. Thepointer parks when the temperature is greater than 1000 °C or less than 0°C. For an invalid brake temperature, the MFD removes the respectivepointer from the display.



5 Test Mode Display

The test mode provides the user with a failure mode annunciation andfamiliarization of the hydraulic system page. The test mode is functionalwhile weight-on-wheels is sensed and the airspeed is less than 50 knots.Pushing and holding the TEST button on the appropriate DC-550 DisplayController initiates the test mode if the hydraulic system page was beingdisplayed. The MFD then displays the test page format as long as the testmode is active. Figure 2-1-50 shows the test mode format for the hydraulicsystem page.

HYDRAULICS

QTY

PRESS

BRAKES

TEMP

IBOB IB OBELEC PUMP

SYS 2

- - - -

- - -

SYS 1

- - - -

- - -HYDRAULICPRESSURE


HYDRAULICELECTRICAL

PUMP ANNUNCIATORS(AMBER DASHES)


Figure 2-1-50. Hydraulic System Page - Test Mode



(c) Takeoff System Page Display

The takeoff system page format contains display information that provides aschematic representation of the aircraft which allows the flight crew to monitorthe status of each door. The page format also provides engine takeoffinformation and digital readouts of the oil levels. White separator lines areshown between the various display information. All legends and labels aredisplayed in white. Figure 2-1-51 shows the display format for the takeoffsystem page.

ENGINE

TAKEOFF MODE:

REF A-ICE: - - -

DOORS

REF TO TEMP: - - -

OIL LVL 7 QT7 QT

- - - - - -

C

SERVICEDOOR 1

FUELINGDOOR

EMERGENCYEXIT DOOR 2

ACCESSDOOR 2

AIRCRAFTSCHEMATIC

BAGGAGEDOOR

EMERGENCYEXIT DOOR 1

MAINDOOR

ACCESSDOOR 1

ELECTRICBAY DOOR

LEGENDS/LABELS

ENGINE 1AND OIL LEVEL

DISPLAY

ENGINETAKEOFF

DATAREADOUTS

ENGINE

REF A-ICE: OFF

DOORS

REF TO TEMP: - 99 C

OIL LVL 1 QT3 QT

T/O-1

DOOROPEN

AIRCRAFT DOOROPEN ANNUNCIATION

ENGINETAKEOFF

DATAREADOUTS

TAKEOFF MODE:

SEPARATORLINES

AIRCRAFTSCHEMATIC

ENGINE TAKEOFFDATA STATUS

ARROW

EXCEEDANCEBOX


Figure 2-1-51. Takeoff System Page Format



1 Engine Takeoff Data

The DAU provides the reference data to the IAC for the engine takeoff datadisplay. This data is then displayed in a digital format. Engine modes,reference temperatures, and anti-ice status are displayed in various statesusing different colors and graphic indications to help the pilot or copilotidentify the aircraft’s current takeoff status. The takeoff mode is displayedin a digital readout as specified in Table 2-1-47.

Table 2-1-47. Takeoff Mode Annunciations

Takeoff Mode 1 Takeoff Mode 2 Annunciation Color

0 0 T/O-1 Table 2-1-48

0 1 ALTTO-1 Table 2-1-48



The reference temperature is displayed in cyan in a digital readout. Thereadout has a range from -99 °C to +99 °C. A white °C label is displayednext to the temperature readout digits. The anti-ice status is alsodisplayed in a digital readout. The anti-ice status is an amber ONannunciation if the bleed position indicates on, and an amber OFFannunciation if it indicates off. If weight on wheels indicates an in airstatus, the takeoff mode, temperature reference, and anti-ice statusreadouts remain blank. If weight on wheels indicates an on ground statusfor more than five seconds, then the logic defined in Table 2-1-48 isimplemented to drive the readouts.



Table 2-1-48. Engine Takeoff Data Logic

Engine Data Readouts Input Signals

Condition

T/OMODE

Display(Table2-1-47)

REF TOTEMP

REFA-ICE

Color TOTempStatus

0

TOTempStatus

1

ToTempStatus

2

Flex/TWRTemp

Takeoff

TempFault

1 TakeoffMode

--- --- W NotSet

NotSet

NotSet

NotSet

NotSet

2 TakeoffMode

TakeoffTemperature

BleedPosition

C Don’tCare

Don’tCare

Don’tCare

Set NotSet

3 TakeoffMode

TakeoffTemperature

BleedPosition

A Set NotSet

NotSet

NotSet

NotSet

4 TakeoffMode

TakeoffTemperature

BleedPosition

A NotSet

Set NotSet

NotSet

NotSet

5 TakeoffMode

TakeoffTemperature

BleedPosition

A Set Set NotSet

NotSet

NotSet

6 TakeoffMode

TakeoffTemperature

BleedPosition

A NotSet

NotSet

Set NotSet

NotSet

7 --- --- --- A Don’tCare

Don’tCare

Don’tCare

Don’tCare

Set

A = AmberC = CyanW = White



A white engine takeoff data status arrow identifies the current takeoffcondition. If condition 3 exists, the arrow is placed next to the takeoffmode readout. If condition 4 exists, the arrow is placed next to the takeofftemperature readout. If condition 5 exists, the arrow is placed next to theanti-ice status readout.

For an invalid takeoff temperature or bleed position, three amber dashes(---) replace the digits in the respective readout. White dashes replace thedigits in all engine takeoff data readouts if the takeoff temperature status isinvalid.

2 Oil Level Display

The DAU provides the oil level data for each engine to the IAC. These oillevels are then displayed in digital readouts. Each readout has a rangefrom 0 to 15 quarts with a resolution of one quart. Each readout has twodisplay colors: amber for oil levels between 0 and 5 quarts, and green(normal) for oil levels between 6 and 15 quarts. When the oil level for anengine exceeds an exceedance limit, an exceedance box is drawn aroundthe digits in the color of the exceedance. A white QT label is displayednext to each digital readout. For an invalid oil level, an amber dashreplaces the digit in the readout and the box around the digit is notdisplayed.

The MFD removes the respective engine oil level readout from the displayif the following conditions exist:

• The N2 value is greater than 50% for one engine and weight on wheelsindicates an on ground status

• Weight on wheels indicates an in air status

The MFD removes the oil level readouts for both engines if the followingconditions exist:

• The N2 value is greater than 50% for both engines and weight onwheels indicates an on ground status

• Weight on wheels indicates an in air status.

3 Door Status Display

A white schematic representation of the aircraft is displayed with coloredboxes to indicate the status of each door. When a door changes status,the color of the representative graphic changes color as defined below:

Door Status Graphic Color

Closed Green

Open Red



The MFD displays a door open message below the aircraft schematic ifany door is in an open status. The message is outlined by a red box. If adoor open status occurs while the takeoff system page is displayed and thedoor message is not already being displayed, the outline box flashes forfive seconds at a rate of one second on/0.5 seconds off, then remainssteady. If a door open status occurs while the system page is notdisplayed and is still present the next time the takeoff system page isselected, the door open message is boxed, flashes for five seconds at arate of one second on/0.5 seconds off, and then remains steady. For eachinvalid door status, the MFD replaces the respective door graphic with anamber X.

4 Test Mode Display

The test mode provides the user with a failure mode annunciation andfamiliarization of the takeoff system page. The test mode is functionalwhile weight-on-wheels is sensed and the airspeed is less than 50 knots.Pushing and holding the TEST button on the appropriate DC-550 DisplayController initiates the test mode if the takeoff system page was beingdisplayed. The MFD then displays the test page format as long as the testmode is active. Figure 2-1-52 shows the test mode format for the takeoffsystem page.

ENGINE

TAKEOFF MODE:

REF A-ICE: - - -

DOORS

REF TO TEMP: - - -

OIL LVL - QT- QT

- - - - - -

C

INVALID OIL LEVELREADOUTS

(AMBER DASHES)


AD-51583@

ENGINETAKEOFF

DATAREADOUTS

(AMBER DASHES)

AIRCRAFT DOORSINVALID(AMBER)

Figure 2-1-52. Takeoff System Page - Test Mode



(d) Environmental Control System Page Display

The environmental control system page format contains display information inthe form of scales and digital readouts which provide identification andseparation of the various environmental control parameters. Headers andheader lines are displayed in white. Figure 2-1-53 shows the display formatfor the environmental control system page.

ECS

CABIN TEMP

CKPT TEMP

-29

-28

C

C

OXY BLEED

PRESS TEMP

PSI1800

HEADERS ANDHEADER LINES

OXYGEN PRESSUREDISPLAY

ENGINEBLEEDTEMPERATUREDISPLAY

CABIN AND COCKPITTEMPERATURE

READOUTS


Figure 2-1-53. Environmental Control System Page Format

1 Cabin Temperature Readout

The DAU provides the cabin temperature data to the IAC. The cabintemperature is then displayed in a digital readout. The readout has arange of -54 °C to +54 °C with a resolution of one degree. The readoutdigits are displayed in green. A white °C label is displayed next to thereadout digits. A white CABIN TEMP label annunciates the readout. Foran invalid cabin temperature, three amber dashes (---) replace the readoutdigits.



2 Cockpit Temperature Readout

The DAU provides the cockpit temperature data to the IAC. The cockpittemperature is then displayed in a digital readout. The readout has arange of -54 °C to +54 °C with a resolution of one degree. The readoutdigits are displayed in green. A white °C label is displayed next to thereadout digits. A white CKPT TEMP label annunciates the readout. For aninvalid cockpit temperature, three amber dashes (---) replace the readoutdigits.

3 Oxygen Pressure Display

The oxygen pressure is displayed as a vertical scale and digital readout.The vertical scale has a range of 0 psi to 2000 psi, with the bottom of thescale representing 0 psi, and the top representing 2000 psi. The verticalscale consists of three color bands and a pointer. The red lower bandrepresents oxygen pressures from 0 psi to 250 psi. The amber middleband represents oxygen pressures from 250 psi to 410 psi. The uppergreen (normal) band represents oxygen pressures from 410 psi to 2000psi.

The pointer moves linearly between the top and bottom of the scale withrespect to the color bands to indicate the oxygen pressure. The pointerreflects the color of the band to which it is pointing. A digital readout ofthe pointer position is displayed below the scale. The readout has a rangeof 0 psi to 2080 psi. A white PSI label is displayed next to the readoutdigits.

The readout digits are displayed in the same color as the pointer. Whenthe oxygen pressure exceeds an exceedance limit, an exceedance box isdrawn around the readout in the color of the exceedance. The pointer alsochanges color accordingly. If the oxygen pressure exceeds the limits ofthe vertical scale, the pointer parks itself at the respective end of the scale.For an invalid oxygen pressure, four amber dashes (----) replace thereadout digits, the exceedance box is not displayed, and the pointer isremoved from the display.

4 Engine Bleed Temperature Display

The MFD displays the left and right cooler outlet temperatures (enginebleed temperatures) as a vertical scale. The vertical scale has a range of-40 °C to 400 °C, with the bottom of the scale representing 180 °C, andthe top representing 350 °C. The vertical scale consists of three colorbands and two pointers. A white lower band represents bleedtemperatures from 0 °C to 260 °C. The green (normal) middle bandrepresents bleed temperatures from 261 °C to 304 °C. An amber upperband represents bleed temperature from 305 °C to 400 °C.



The pointers move linearly between the top and bottom of the scale withrespect to the color bands to indicate the cooler outlet temperatures. Forvalues between 350 °C and 400 °C, or between -40 °C and 180 °C, thepointer remains parked at the top and bottom of the scale respectively.The pointer on the left side of the scale represents the left cooler outlettemperature. The pointer on the right side of the scale represents the rightcooler outlet temperature.

Each pointer reflects the color of the band to which it is pointing in theupper region of the scale. In the lower region of the scale, the pointercolor is as follows:

• The left and right pointers are green for normal operation

• The left pointer is amber when the "BLD 1 TEMP" CAS message isdisplayed

• The left pointer is amber when the "BLD 1 TEMP" CAS message isdisplayed.

When a cooler outlet temperature exceeds the limits of a color band, thepointer changes to the new exceedance color. For an invalid cooler outlettemperature, the MFD removes the respective pointer from the display.



5 Test Mode Display

The test mode provides the user with a failure mode annunciation andfamiliarization of the environmental control system page. The test mode isfunctional while weight-on-wheels is sensed and the airspeed is less than50 knots. Pushing and holding the TEST button on the appropriateDC-550 Display Controller initiates the test mode if the environmentalcontrol system page was being displayed. The MFD then displays the testpage format as long as the test mode is active. Figure 2-1-54 shows thetest mode format for the environmental control system page.

ECS

CABIN TEMP

CKPT TEMP

- - -

- - -

C

C

OXY BLEED

PRESS TEMP

PSI- - - -

OXYGEN PRESSUREREADOUT

(AMBER DASHES)

CABIN AND COCKPITTEMPERATURE

READOUTS(AMBER DASHES)


AD-51585@

Figure 2-1-54. Environmental Control System Page - Test Mode



(e) Fuel System Page Display

The fuel system page format contains display information that provides aschematic representation of the fuel system for the pilot or copilot. Digitalreadouts are provided for the total fuel quantity, total fuel used, and fuel tanktemperature. Header lines and static tank lines are displayed in white. Figure2-1-51 shows the display format for the takeoff system page.

FUEL

TOTAL

USED

LB

LB

TANK 2

CTEMP

1350

-35

TANK 1

PUMPOFF

LB650

PUMPA C

LB700

TOTAL FUELQUANTITYREADOUT

HEADER ANDSTATIC TANK

LINES

TANK 2 FUELQUANTITYREADOUT


TANK 1 FUELQUANTITY

SCALE

TANK 2 FUELQUANTITYSCALE

TANK 1 FUELPUMP STATUSANNUNCIATOR


FUEL USEDREADOUT

FUEL TANKTEMPERATURE

READOUT


350

AD-51586@

Figure 2-1-55. Fuel System Page Format

1 Fuel Tank Quantity Display

The DAU provides the tank 1 and tank 2 fuel guantity in pounds to the IAC.The IAC converts pounds to kilograms if the English/Metric strap indicatesmetric. Each fuel tank quantity is then displayed as a vertical scale anddigital readout.



The vertical scale for the tank 1 fuel quantity display is positioned alongthe left edge of the display. The vertical scale for the tank 2 fuel quantitydisplay is positioned along the right edge of the display. Each scale has arange from 0 LB (0 KG) to 5000 LB (2270 KG). Each scale contains threecolor bands which represent the following fuel quantities:

• Red band (lower) represents fuel quantities from 0 LB (O KG) to 620LB (280 KG)

• Amber band (middle) represents fuel quantities from 630 LB (290 KG)to 880 LB (400 KG)

• Green (normal) band (upper) represents fuel quantities from 890 LB(410 KG) to 5000 LB (2270 KG).

A pointer moves linearly between the top and bottom of each scale withrespect to the color bands to indicate the fuel tank quantity. Each pointerreflects the color of the band to which it is pointing.

A digital readout of the pointer position is displayed next to each verticalscale. Each readout has a range of 0 to 5000 pounds with a resolution of10 pounds when english units are selected. The readout range is from 0to 3000 kilograms with a resolution of 10 kilograms when metric units areselected. A white LB or KG label is displayed next to the readout digits,depending on whether pounds or kilograms is selected for display. Thereadout digits are displayed in the same color as the respective pointer.

The readouts are not boxed for normal conditions. When a digital fuelquantity (excluding total quantity) exceeds an exceedance limit, anexceedance box is shown around the readout digits in the color of theexceedance. When the fuel quantity exceeds the limits of a particular colorband on the vertical scale, the pointer changes color accordingly. If thefuel quantity exceeds the limits of the scale, the pointer parks itself at therespective end of the scale. For an invalid fuel tank quantity, three amberdashes (---) replace the respective readout digits, the exceedance box isnot displayed, and the pointer is removed from the display.

2 Fuel Tank Pump Status Annunciator

The DAU provides the status of each fuel tank pump to the IAC. Eachstatus is considered individually unless all pumps indicate an off status.Each fuel status is displayed as a three character status field with possiblevalues of A, B, C, or OFF located in the left, center, and right positionsrespectively. A white PUMP label is displayed above each pump statusfield.



A green A annunciation is displayed in the left position if power contactor Afor that fuel pump is active. The position is blank if power contactor A isinactive. A green B annunciation is displayed in the center position ifpower contactor B for that fuel pump is active. The position is blank ifpower contactor B is inactive. A green C annunciation is displayed in theright position if power contactor C for that fuel pump is active. Theposition is blank if power contactor C is inactive. If all power contactors forthat fuel pump are inactive, a green OFF annunciation is displayed in thepump’s status field. For an invalid fuel tank pump status, an amber dash(-) is placed in the corresponding pump status field position.

3 Total Fuel Quantity Readout

The total fuel quantity is displayed in a digital readout. The readout has arange from 0 to 9990 pounds with a resolution of 10 pounds when englishunits are selected. The readout has a range from 0 to 6000 kilograms witha resolution of 10 kilograms when metric units are selected. A white LB orKG label is displayed next to the readout digits, depending on whetherpounds or kilograms is selected for display. The readout digits aredisplayed in green for normal conditions.

The readouts are not boxed for normal conditions. If either tank 1 or tank2 fuel quantities enter into an exceedance range, an exceedance box isshown around the readout digits in the same color as the greaterexceedance. For an invalid total fuel quantity, four amber dashes (----)replace the readout digits.

4 Fuel Used Readout

The DAU provides the IAC with the fuel flow data from each tank. Thisdata is then integrated and added to a running total to obtain the fuel usedquantity which is displayed in a digital readout. The IAC converts the fuelused value from pounds to kilograms if the English/Metric strap indicatesMetric. An IAC reset effects the fuel used readout based on whether thereset occurred while the aircraft was on the ground or in the air.

The readout has a range from 0 to 9990 pounds with a resolution of 10pounds when english units are selected. The readout has a range from 0to 4530 kilograms with a resolution of 10 kilograms when metric units areselected. A white LB or KG label is displayed next to the readout digits,depending on whether pounds or kilograms is selected for display. Thereadout digits are displayed in green. For an invalid fuel used value, fouramber dashes (----) replace the readout digits.



5 Fuel Tank Temperature Readout

The DAU provides the fuel tank temperatures to the IAC. The fuel tanktemperature is then displayed in a digital readout. The readout has arange from -64 °C to +64 °C with a resolution of one degree. Fortemperatures less than 0 °C, a minus sign (-) is displayed with the digits.The readout digits are displayed in green (normal) for fuel tanktemperatures between -39 °C to +64 °C. The readout digits are displayedin amber when the fuel tank temperature drops below -39 °C. A white °Clabel is displayed next to the readout digits.

When the fuel tank temperature exceeds an exceedance range, anexceedance box is shown around the readout digits in the same color asthe exceedance. For an invalid fuel tank temperature, three amber dashes(---) replace the readout digits.

6 Test Mode Display

The test mode provides the user with a failure mode annunciation andfamiliarization of the fuel system page. The test mode is functional whileweight-on-wheels is sensed and the airspeed is less than 50 knots.Pushing and holding the TEST button on the appropriate DC-550 DisplayController initiates the test mode if the fuel system page was beingdisplayed. The MFD then displays the test page format as long as the testmode is active. Figure 2-1-56 shows the test mode format for the fuelsystem page.

FUEL

TOTAL

USED

LB

LB

TANK 2

CTEMP

- - - -

- - -

TANK 1

PUMP- - -

LB- - - -

PUMP- - -

LB- - - -

TOTAL FUELQUANTITYREADOUT

(AMBER DASHES)


(AMBER DASHES)


(AMBER DASHES)


(AMBER DASHES)


(AMBER DASHES)

FUEL USEDREADOUT

(AMBER DASHES)

FUEL TANKTEMPERATURE



- - -

AD-51587@

Figure 2-1-56. Fuel System Page - Test Mode



(11) MFD Test Mode Display

The system test mode provides the pilot or copilot with a failure mode annunciationand familiarization of the MFD format. The test mode is functional whileweight-on-wheels is sensed and the airspeed is less than 50 knots. Pushing andholding the TEST button on the appropriate DC-550 Display Controller initiates thetest mode for the corresponding MFD. The MFD then displays the test page formatas long as the test mode is active. Figure 2-1-57 shows the test page format for theMFD.

When the MFD test mode is selected, the currently selected system page isdisplayed in its test page format. If the plan mode is the current display mode, themap mode is forced on when the test mode is initiated. The plan mode is re-enabledwhen the TEST button is released.

The MFD displays the following items in their invalid state:

• Bezel menu

• Distance

• Map mode heading display

• SAT

• Time-to-go

• TAS

• TAT

• Weather radar

The following display items are forced off the display:

• Airports and airport identifiers

• Checklist display

• Designator information

• Drift bug

• Flight plan data (waypoints, identifiers, track lines)

• Heading select bug

• Lateral deviation

• NAVAIDS and NAVAID identifiers

• TCAS display

• Wind display

• WX video data



INVALID SAT (AMBER)

INVALID DISTANCE(AMBER)

AD-51241@

INVALID TAT (AMBER)

INVALID TAS (AMBER)

HEADING TESTANNUNCIATOR

(RED)

INVALID TTG(AMBER)

INVALIDCURRENTHEADING

N

S

33

15

3012

WE

246

21

3

- - -- - -- - -

SATTATTAS

FMS

- - -- - -

- - -

50 50

MENU INOP

HDGTEST


WX

FAILED WX(AMBER)

Figure 2-1-57. MFD Test Mode Display



K. Engine Instrument Crew Alerting System (EICAS) Display Interface

NOTE: See Figure 2-1-58 (Sheets 1 and 2) for a block diagram of the EICASinterface.

The EICAS display utilizes four primary components as follows:

• DU-870 Display Unit No. 3

• BL-871-851 EICAS Bezel Controller (1)

• DC-550 Display Controller (1)

• DA-800 Data Acquisition Units (DAU) (2)

• Symbol Generator (part of the IC-600 IAC) (2)

The DAU is the central data collection point for the EICAS. Left side aircraft and enginesensors connect to DAU No.1. Right side aircraft and engine sensors connect to DAU 2.

The DAUs receive discrete, digital and analog data from various aircraft instruments andcomponents either directly, or from each IAC. This data includes engine data from the FullAuthority Digital Engine Computer (FADEC). The FADEC data is digitized (if necessary)and sent to the symbol generator in each IAC.

The controlling IAC sends EICAS data to the EICAS display unit for display. An ARINC429 wraparound signal from the EICAS display unit is sent to both IACs where it iscompared with the original sensor data in the IAC to ensure validity of the displayed data.The dimming panel controls the brightness of the EICAS display.

Each DAU consists of two independent and isolated channels with identical hardware andsoftware. Each channel contains its own power supply for connection to an aircraftcircuit breaker. Both channels perform the same computations on identical input data.Both channels are continuously monitored by the IAC.

The IAC displays errors or detected differences on the EICAS parameters between the twochannels as messages in the CAS section of the EICAS display. The control knob on theEICAS display bezel allows the pilot or copilot to scroll through the CAS messages. Themaster warning and master caution annunciators on the glareshield are used toacknowledge warning and caution messages shown on the CAS display.

If the EICAS display unit fails, the DU revision switch on either reversionary panel can beused to move the EICAS display format to an MFD unit. The control knob on the MFDbezel can then be used scroll through the CAS messages.



(Blank Page)



10" MAX STUB

11413

24

23

(H)(W)(L)

LIGHTING COMMON

0-5 VAC EDGE LIGHTING 0-5 VAC

BRIGHTNESS POT

28 V DCINPUT POWER

POWER GROUND

28 V DCBUS 1

212563

(B0)(B1)

SET KNOB242526 SET KNOB COMMON

(H)(L)

SET KNOB3435

HL

DC/IC BUSOUT

DC-550 DISPLAYCONTROLLER NO. 1

242526 SET KNOB COMMON

(H)(L)

SET KNOB

87H

L

DC-550 DISPLAYCONTROLLER NO. 2

DU-870 NO. 3 (EICAS)

101102103

104105106

132J1

115J1115J1

C115J1 C115J1

AD-51244,SH1@

IC-600 NO.1

1011

+

-DU HDLC OUT

1617

1415

+

-

+

-

DU HDLC NO. 3 IN(UNTERMINATED)

DU HDLC NO. 1 IN(TERMINATED)

45

46

47

+

-DU HDLC NO. 2 IN

OPT HDLC NO. 2 TERM

EICAS DW ESS DC

IC-600 NO.2

1011

+

-DU HDLC OUT

1617

1415

+

-

+

-

DU HDLC NO. 3 IN(UNTERMINATED)

DU HDLC NO. 1 IN(TERMINATED)

4546

47

+

-DU HDLC NO. 2 IN

OPT HDLC NO. 2 TERM

AIRCRAFT DIMMINGCONTROL

DC/IC BUSOUT

23" MIN

16" MAX STUB

23" MIN

16" MAX STUB

23" MIN

16" MAX STUB

BEZEL COMMON3839

3536

3132

23" MIN

10" MAX STUB

TO 130J1-35,36AND 131J1-35,-36.

TO 130J1-38,-39AND 131J1-38,-39.

HDLC NO. 1 IN

HDLC NO. 2 IN

HDLC OUT

TO C130J1-38, 39C131J1-38, 39

132J1

190J2A

190J2B

C190J2B

C190J2A

DU NO.3 - EICAS

15

16

H

L

DC/ICBUS IN

190J2A

15

16

H

L

DC/ICBUS IN

C190J2A

TO C130J1 -35, 36C131J1-35, 36

Figure 2-1-58 (Sheet 1). EICAS Interface Diagram



4647

4445

4849

6869

HL

HL

HL

HL

L-FADEC CH AARINC INPUT

SHIP'S CLOCKARINC INPUT

FUEL COMPUTERARINC INPUT

FLIGHT DATAREC OUT

4445

4243

HL

HL

L-FADEC CH BARINC INPUT

STALL WARN CH. AARINC INPUT

104105106

104105106

136J2B

136J1B

136J1/J2

TXRX

GND

RS-232CHANNEL A

TXRX

GND

RS-232CHANNEL B

RS-232 PILOT'STEST CONNECTOR

A/CTRANSMITTER

136J1B

136J2B

A/CTRANSMITTER

NO. 1FADEC

136J1A

DA-800 DAU NO. 1

104105106

RS-232 COPILOT'STEST CONNECTOR

104105106

TXRX

GND

RS-232 COPILOT'STEST CONNECTOR

RS-232CHANNEL A

TXRX

GND

RS-232CHANNEL B

VARIOUS DISCRETE ANALOGINPUTS FOR EICASMESSAGES/DISPLAYS

DA-800 DAU NO. 2DISCRETE/ANALOGINPUTS/OUTPUTS:

REFER TO SECTION 3, INTERCONNECT

INFORMATIONFOR SIGNAL NAME./

CONNECTOR PIN DATA.

DISCRETE/ANALOGINPUTS/OUTPUTS:REFER TO SECTION 3,INTERCONNECTINFORMATIONFOR SIGNAL NAME./CONNECTOR PIN DATA.

VARIOUS DISCRETE ANALOGINPUTS FOR EICAS

MESSAGES/DISPLAYS

137J1/J2

137J1B

137J2B

137J1B

RS-232 PILOT'STEST CONNECTOR

4041

HL

MAINT COMPARINC OUTPUT

MAINTENANCECOMPUTER

99100

103104

105106

10169

SIG GND NO. 1 CH ASIG GND NO. 2 CH A

28V DC NO. 1 CH A28V DC NO. 2 CH A

PWR GND NO. 1 CH APWR GND NO. 2 CH A

CHASSIS GND CH AID NO. 3 CH A

100

103104

105106101

69

SIG GND NO. 2 CH B

28V DC NO. 1 CH B28V DC NO. 2 CH B

CHASSIS GND CH BID NO. 3 CH B

99 SIG GND NO. 1 CH B

136J2A

PWR GND NO. 1 CH BPWR GND NO. 2 CH B

80 W.O.W.

9899

100101

HL

HL

RS-422 XMTCH A

RS-422 RCVCH A

190J2

5051

IC-600 NO. 1 IC-600 NO. 2C190J2

A B A B

8182

3132

4243

ARINC 429XMTR NO. 2A

H

L

HL

IAC NO. 1ARINC INPUT

5051

HL

ARINC 429XMTR NO. 2A

5051

ARINC 429XMTR NO. 2B

HL

3132

8182

3132

4142

3738

4142

3132

HL

DAU NO. 1 CH A

DAU NO. 2CH A

429 OUTTO DAU

DAU NO. 2CH B

DAU NO. 1 CH B

DAU NO. 1CH A

DAU NO. 2CH A

DAU NO. 1CH B

DAU NO. 2CH B

GROUND =AIRCRAFT ON GROUND

28V5051

HL

ARINC 429XMTR NO. 2B

100

103104

105106

101

68

SIG GND NO. 2 CH B

28V DC NO. 1 CH B28V DC NO. 2 CH B

CHASSIS GND CH BID NO. 2 CH B

99SIG GND NO. 1 CH B

PWR GND NO. 1 CH BPWR GND NO. 2 CH B

80W.O.W.

137J1A

99100

103104

105106

10168

SIG GND NO. 1 CH ASIG GND NO. 2 CH A

28V DC NO. 1 CH A28V DC NO. 2 CH A

PWR GND NO. 1 CH APWR GND NO. 2 CH A

CHASSIS GND CH AID NO. 2 CH A

137J2A

4647

4445

4849

HL

HL

HL

R-FADEC CH AARINC INPUT

CABIN PRESS.SENSOR UNIT

FUEL COMPUTERARINC INPUT

A/CTRANSMITTER

137J2B

4445

4243

HL

HL

R-FADEC CH BARINC INPUT

STALL WARN CH. AARINC INPUT

A/CTRANSMITTER

NO. 2FADEC

4041

HL

MAINT COMPARINC OUTPUT

MAINTENANCECOMPUTER

HL

HL

RS-422 XMTCH A

RS-422 RCVCH A

9899

100101

HL

HL

HL

HL

HL

HL

HL

HL


AD-51244-R1,SH2@

DAU 1BDC BUS1

28VDAU 1AESS DC BUS1

28VDAU 2A ESS DC BUS1

28VDAU 2B DC BUS1

Figure 2-1-58 (Sheet 2). EICAS Interface Diagram



L. EICAS Display Operation

The EICAS display format provides engine information to the pilot. The EICAS display alsoprovides a crew alerting system (CAS), landing gear information, cabin environment andtrim data. The EICAS display format consists of five major display areas as follows:

• Engine Instrument Section

• Cabin and APU Section

• CAS Message Section

• Flight Control Information Section

Figure 2-1-59 shows the display format for EICAS information.

If an exceedance occurs on the EICAS, it causes the box around the digital readout toflash for five seconds at a rate of one second on and 0.5 seconds off, then remain steady.If a parameter is currently in exceedance and it enters a higher level exceedance, the boxflashes for five seconds and then remains steady. The box changes to the exceedancecolor but does not flash if a parameter currently in exceedance enters a lower levelexceedance.

If a parameter is currently in exceedance with its box flashing and it enters a higher levelexceedance, the box changes to the new exceedance color and flashes for the remainderof the lower level exceedance’s five second duration. If a parameter is currently inexceedance with its box flashing and it enters a lower level exceedance, the box changesto the new exceedance color and flashes for the remainder of the higher levelexceedance’s five second duration. The box is immediately removed if a parameterbecomes invalid while in an exceedance with the box flashing. If a parameter is invalidand then becomes valid within an exceedance range, the box and digits change to theexceedance color and the box flashes for five seconds and then remains steady.



CASMESSAGESECTION

ENGINEINSTRUMENT

SECTION

UPPER FLIGHTCONTROL

INFORMATIONSECTION

CABIN/APUSECTION

LOWER FLIGHTCONTROL

INFORMATIONSECTION

AD-51245@

Figure 2-1-59 (Sheet 1). EICAS Display Format



FLAPS

SPLRS

0

CLD

LDG GEAR

CAB ALT

CAB P

CAB RATE

APU

ROLL

YAW

PITCH

UP

9

88.1 T/0-1

UP UP UP

OIL VIB

79 79 100100

PRESS TEMP LP HP

AA

N2

FF

FQ

99.9%

990 PPH

1000 LB

99.9%

990 PPH

1000 LB

ITT

N1

550

75.0

550

75.0ATTCS

88.1

REV REV

IGN IGNA A

END

FT

PSI

FPM

C

7000

7.5

-5000

100% 600

ENGINE FAN SPEED DISPLAY

N1 REQUEST BUG

INTER TURBINETEMPERATURE

DISPLAY

ENGINE TURBINESPEED READOUT

FUEL FLOWREADOUT

FUEL TANKQUANITY

READOUT

OIL PRESSUREDISPLAY

OIL TEMPERATUREDISPLAY

APU TURBINE SPEEDDISPLAY

APU TEMPERATUREDISPLAY

LOW PRESSUREVIBRATION

VERTICAL SCALE

HIGH PRESSUREVIBRATION

VERTICAL SCALE

GROUNDSPOILERDISPLAY

FLAP POSITIONDISPLAY

LANDING GEARDISPLAY

CREW ALERTINGSYSTEMMESSAGE FIELD

"END" STATUSMESSAGE

N1 TARGET BUG ANDDIGITAL READOUT

AUTOMATIC TAKEOFF THRUST CONTROLSYSTEM ENABLED ANNUNCIATOR

AD-51246@

SEPARATORLINES

LEGENDS


CABIN ALTITUDEDISPLAY

ENGINE MODEANNUNCIATOR




FLAPS

SPLRS

22

OPN

LDG GEAR

CAB ALT

CAB P

CAB RATE

APU

ROLL

YAW

PITCH

UP

9

88.1 R-MODE

DN UP

OIL VIB

79 79 100145

PRESS TEMP LP HP

AA

N2

FF

FQ

99.9%

990 PPH

1000 LB

99.9%

990 PPH

700 LB

ITT

N1

550

75.0

550

75.0

88.1

REV REV

IGN IGNA A

FT

PSI

FPM

10000

7.5

-5000

OFF

THRUST REVERSERANNUCIATOR

FADEC IN CONTROLCHANNEL

ANNUNCIATOR

IGNITIONANNUNCIATOR

CABIN DIFFERENTIALPRESSURE DISPLAY

YAW TRIM DISPLAY ROLL TRIM DISPLAY

PITCH TRIMDISPLAY

AILERON MISTRIMARROW

MESSAGESSTATUSLINE

WARNINGMESSAGES(RED)

N1 EXCEDENCE ARC

AD-51247@

ADVISORYMESSAGES(CYAN)

CAUTIONMESSAGES(AMBER)

BATT 1 OVTEMP

BLD 1 LEAK

MAIN DOOR OPN

GPWS INOP

FUEL XFEED FAIL

PITOT 3 INOP

LG/LEVER DISAGREE

BRAKE INBD INOP

DFDR FAIL

ENG1 T/R DISAGREE

E1 FUEL IMP BYP

BRAKE DEGRADED

FUEL XFEED OPEN

HYD2 LO QTY

0 MESSAGES 2


CABIN ALTITUDERATE DISPLAY

APU STATUSANNUNCIATOR




(1) Engine Instrument Section

The engine instrument section of the EICAS display contains engine information forthe following parameters:

• Engine Fan Speed (N1)

• Inter Turbine Temperature (ITT)

• Engine Turbine Speed (N2)

• Fuel Flow

• Fuel Quantity

• Oil Pressure

• Oil Temperature

• Engine Vibration

• Engine Mode

• Ignition Status

• FADEC in Control

• Thrust Reversers

Engine instrument legends N1, ITT, N2, FF, and FQ are displayed as long as thedisplay unit is in EICAS mode. These legends are centered top to bottom in theengine instrument section, dividing the left engine data from the right engine data.The OIL and VIB subsection labels are constantly displayed, along with the PRESS,TEMP, LB and HP indicator labels. The %, PPH, KPH, LB, and KG legends are alsoconstantly displayed. All legends and labels are displayed in white in the engineinstrument section. White separator lines separate the oil and vibration subsectionfrom the main engine instrument section.

(a) Engine Fan Speed (N1) Display

Engine fan speed (N1) data is displayed as an indicator dial and digitalreadout for the left and right engines. Left engine N1 data is shown on theleft, and right engine N1 data is shown on the right. Each engine is controlledby a full authority digital engine computer (FADEC). When the FADECcontrols the engine speed, the N1 request bug is displayed on the outside ofthe indicator dial. The digital readout is displayed inside the indicator dial forthe N1 reference value.



1 Indicator Dial

The indicator dial consists of the following elements:

• Dial scale consisting of two arcs: normal and exceedance

• Dial pointer

• Digital readout

The normal arc is a 200° arc beginning at the 70° compass position andextending in a clockwise direction. The normal arc is scaled as twodegrees of arc per 1% of N1, with the 70° compass position representing0% N1 and the 270° compass position representing 100% N1. Long whitetick marks are placed on the arc at every 40 degrees of compass arcstarting at the 70° compass position. Short white tick marks are placed onthe arc at every 40 degrees of compass arc starting at the 90° compassposition. A double stroked tick mark extending to the right of the normalarc is placed at the 100% N1 position. The tick marks and percent labelsare displayed in white.

The exceedance arc is a 20° arc beginning at the 270° compass positionand extending in a clockwise direction. The exceedance arc is scaled astwo degrees of arc per 1% of N1, with the 270° compass positionrepresenting 100% N1 and the 290° compass position representing 110%N1. The exceedance arc is visible only when the dial pointer and digitalreadout exceed the upper limits of the normal arc range.

The dial pointer, indicating N1, sweeps the normal and exceedance arcswith a resolution of 0.1%. The pointer is limited at 110% for N1 valuesgreater than 110%, and less than or equal to the upper digital readoutlimit. The digital readout has a range of 0.0 % to 199.9% N1 with aresolution of 0.1%. A white box surrounds the digital readout. This boxchanges to red if an exceedance occurs. For invalid N1, four amberdashes (---.-) replace the readout digits and the pointer and exceedancearc are removed.

The N1 indicator dial parameters are defined in Table 2-1-49.

Table 2-1-49. N1 Indicator Dial Parameters

Range Duration ScaleColor

Pointer/ReadoutColor

0.0% to100% of N1

Always White Green

100.1% to110% of N1

When N1exceeds

100%

Red Red



2 N1 Request Bug

The N1 request bug is displayed as a green, truncated triangle and isprovided for both N1 displays. The N1 request bug moves along theoutside edge of the N1 indicator dial. For an invalid N1 request, therespective N1 request bug is removed from the display.

3 N1 Target Bug

The N1 target bug consists of a T symbol and a digital readout and isprovided for both N1 displays. The N1 target moves along the outsideedge of the N1 indicator dial. The N1 target bug digital values aredisplayed in a digital readout. The readout has a range of 05 to 100% witha resolution of 0.1 % RPM. Both the target bug and readout digits aredisplayed in cyan. For an invalid engine target N1, the respective N1target bug is removed and four amber dashes (---.-) replace the readoutdigits.

(b) Inter Turbine Temperature Display

Inter turbine temperature (ITT) data is displayed as an indicator dial and digitalreadout for the left and right engines. Left engine ITT data is shown on theleft, and right ITT engine data is shown on the right. The digital readout isdisplayed inside the indicator dial for the ITT reference value.

The indicator dial consists of the following elements:

• Dial scale consisting of two arcs: normal and exceedance

• Dial pointer

• Digital readout

The normal arc begins at the 70° compass and extends in a clockwisedirection. The normal arc consists of two linear arc segments. The first arcsegment is static and scaled such that 4.33 degrees of arc equals one degreeof change in ITT. The second arc segment is scaled such that 2.49° of arcequals a 1° change in ITT. The second arc segment extends clockwise fromthe end of the static arc segment to the start of the red exceedance arc (redtick mark) as follows:

• When conditions in Table 2-1-50 indicate the red exceedance arc startingpoint is 922 °C, the second arc segment extends clockwise from the end ofthe static arc segment to the red tick mark as shown in Figure 2-1-60.

• When conditions in Table 2-1-50 indicate an engine start has beenrequested and the engine is not running, the second arc segment extendsclockwise from the end of the static arc segment to the red tick mark asshown in Figure 2-1-61.



• When conditions in Table 2-1-50 indicate the red exceedance arc staringpoint is 922 °C and the engine mode is CON, CLB, or CRZ, the second arcsegment extends from the end of the static arc segment to 869 °C, and anamber exceedance arc is drawn from 869 °C to the red tick mark as shownin Figure 2-1-62.

Table 2-1-50. ITT Engine Start Logic

Engine StartRequested

EngineRunning

Red Exceedance ArcStarting Point (Tick Mark)

NotRequested

NotRunning

922°

NotRequested

Running 922°

Requested NotRunning

801°

Requested Running 922°

590 ºC

300 ºC

133º67º922 ºC

ITT Arc

AD-51381@

EXCEEDANCE ARCFROM 922° TO 972°

RED TICKMARK

STATIC ARCSEGMENT

SECOND ARCSEGMENT

Figure 2-1-60. ITT Arc Default



590 ºC

300 ºC

133º67º

801 ºC

AD-51588@

EXCEEDANCE ARCFROM 801° TO 851°

RED TOCKMARK

STATIC ARCSEGMENT

SECOND ARCSEGMENT

Figure 2-1-61. ITT Arc During Engine Start

869 ºC

300 ºC

922 ºC

AD-51589@

EXCEEDANCE ARCFROM 922° TO 972°RED TICK

MARK

STATIC ARCSEGMENT

SECOND ARCSEGMENT

AMBER REGION OFARC FROM

869° TO 922°

AMBER TICKMARK

Figure 2-1-62. ITT Arc With Engine Mode



A red temperature exceedance arc is displayed when the ITT digital readoutand pointer are greater than or equal to the starting point of the redexceedance arc. The red temperature exceedance arc is a 20 degree radialarc (50 °C) beginning at the red exceedance arc starting point and extendingin a clockwise direction.

The dial pointer, indicating temperature, sweeps the normal and exceedancetemperature arcs with a resolution of one degree. The pointer is limited at 300degrees for ITT input values less than or equal to 300 °C. For ITT inputvalues which exceed the range of the exceedance region, the pointer is limitedat the end of the 50 °C temperature exceedance arc.

The digital readout has a range from -65 °C to 1999°C with a resolution of onedegree. The pointer and readout digits are displayed in green for normal ITTconditions. When an exceedance occurs, the pointer and readout digits aredisplayed in the exceedance color.

A white box surrounds the digital readout. When an exceedance occurs, thebox transitions to an amber or red exceedance color. The box also flashes forfive seconds at a rate of one second on and 0.5 seconds off during thetransition to an exceedance color, or from an amber exceedance color to a redexceedance color.

For an invalid ITT, four amber dashes (----) replace the readout digits, the boxis displayed in its normal state, and the pointer and exceedance arc areremoved. For an invalid engine start request, engine running, and enginemode, the ITT display defaults to the display shown in Figure 2-1-60.

(c) Engine Turbine Speed (N2) Display

Engine turbine speed (N2) data is displayed in a digital readout for the left andright engines. The left engine N2 data is shown on the left, and right engineN2 data is shown on the right. The digital readout has a range from 0.0% to199.9% with a resolution of 0.1 percent. A percent (%) symbol is displayednext to the readout digits. The readout digits are displayed in green fornormal N2 conditions.

The digital readout is not boxed during normal conditions. When an N2 valueenters an exceedance range, a red box is shown around the readout digitswhich are also displayed in red. For an invalid N2, four amber dashes (---.-)replace the readout digits.



(d) Fuel Flow Readout

Fuel flow is measured in pounds per hour (PPH). The DAU provides the fuelflow in PPH to the IC-600 IAC. The IAC converts the value to kilograms perhour (KPH) if the English/Metric strap indicates metric. The EICAS displayannunciates which system is being used with a white PPH or KPH label.

Fuel flow (FF) for the left and right engines is displayed in a digital readoutabove the fuel tank quantity data. Fuel flow data for the left engine is shownon the left, and fuel flow data for the right engine is shown on the right. Thereadout has a range from 0 PPH to 4000 PPH with a resolution of 10 PPH forenglish units. The readout has a range from 0 KPH to 1820 KPH with aresolution of 10 KPH for metric units. The readout digits are displayed ingreen. For an invalid fuel flow, four amber dashes (----) replace the readoutdigits.

(e) Fuel Tank Quantity Readout

Fuel tank quantity is measured in pounds (LB). The DAU provides the IACwith fuel tank quantity in pounds. The IAC converts the value to kilograms(KG) if the English/Metric strap indicates metric. The EICAS displayannunciates which system is being used with a white LB or KG label.

Fuel tank quantity (FQ) for the left wing and right wing fuel tanks is displayedin a digital readout below the fuel flow data. The left wing fuel tank quantity isshown on the left, and the right wing fuel tank quantity is shown on the right.The readout has a range from 0 LB to 5000 LB with a resolution of 10 poundsfor english units. The readout has a range from 0 KG to 3000 KG with aresolution of 10 kilograms for metric units.

The readout digits change color as function of the remaining fuel tank quantity.readout digits are displayed in green for a full fuel tank quantity (5000 LB or3000 KG). Readout digits are displayed in amber when the fuel tank quantityreaches 880 LB (400 KG). Readout digits are displayed in red for a fuel tankquantity less than 620 LB (280 KG). When a fuel tank quantity exceeds alimit, an exceedance box in the color of the exceedance is shown around thereadout digits.

If the primary fuel tank quantity is invalid, the secondary fuel tank quantity islatched as the only data source and remains latched until the IAC isre-powered. If the secondary fuel tank quantity is invalid, four amber dashes(----) replace the readout digits and the exceedance box is removed.



(f) Oil Pressure Display

Engine oil pressure is measured in pounds per square inch (PSIg). Oilpressure data is displayed as a vertical scale and a digital readout for the leftand right engines. Left engine oil pressure is shown on the left, and rightengine oil pressure is shown on the right.

The vertical has a range from 0 PSig to 99 PSIg, with the bottom of the scalerepresenting 0 PSIg and the top representing 99 PSIg. The vertical scalecontains three color bands as follows when N2 is less than 88%:

• The red band (15% of scale length) represents oil pressures from 0 PSIgto 31 PSIg

• The green band (70% of scale length) represents oil pressures from 32PSIg to 90 PSIg

• The amber band (15% of scale length) represents oil pressures from 91PSIg to 99 PSIg.

When N2 is greater than 88%, the vertical scale displays an additional amberband as follows.

• The red band (15% of scale length) represents oil pressures from 0 PSIgto 31 PSIg

• The lower amber band (15% of scale length) represents oil pressures from32 PSIg to 49 PSIg.

• The green band (50% of scale length) represents oil pressures from 50PSIg to 90 PSIg

• The upper amber band (15% of scale length) represents oil pressures from91 PSIg to 99 PSIg.

Two pointers are positioned on the vertical scale; one on the left side for theleft engine oil pressure, and the other on the right side for the right engine oilpressure. Each pointer moves in linearly between the top and bottom of thescale with respect to the three color bands. When the oil pressure exceedsthe defined limits for a specific band, the pointer changes color. If the oilpressure exceeds the scale limits, the pointer parks itself at the respective endof the scale.

A digital readout of the pointer position is provided for each engine oilpressure. The readout has a range from 0 PSIg to 150 PSIg with a resolutionof one PSIg. The readout digits are displayed in the same color as thecorresponding pointer on the vertical scale. When a digital engine oil pressurevalue exceeds a limit, the readout digits are boxed in the respective color ofthe exceedance. For an invalid engine oil pressure, two amber dashes (--)replace the readout digits and the pointer and exceedance box are removed.



(g) Oil Temperature Display

Engine oil temperature is measured in degrees Celsius. The oil temperaturefor the left and right engines is displayed as a vertical scale and a digitalreadout. Left engine oil temperature is shown on the left, and right engine oiltemperature is shown on the right.

The scales have a range from 0 °C to 180 °C, with the bottom of the scalerepresenting 0 °C and the top representing 180 °C. The vertical scalecontains three bands as follows:

• The amber band represents oil temperatures from 0 °C to 21 °C

• The green band (normal) represents oil temperatures from 21 °C to 127 °C

• The red band represents oil temperatures from 127 °C to 180 °C.

Two pointers are positioned on the vertical scale; one on the left side for theleft engine oil temperature, and the other on the right side for the right engineoil temperature. Each pointer moves linearly between the top and bottom ofthe scale with respect to the three color bands. When the engine oiltemperature exceeds the defined limits for a specific band, the pointerchanges color. If the engine oil temperature exceeds the scale limits, thepointer parks at the respective end of the scale.

A digital readout of the pointer position is provided for each engine oiltemperature. The readout has a range from -80 °C to 300 °C with a resolutionof one degree. The readout digits are displayed in the same color as thecorresponding pointer on the vertical scale. When a digital engine oiltemperature value exceeds a limit, the readout digits are boxed in therespective color of the exceedance. For an invalid engine oil temperature, twoamber dashes (--) replace the readout digits and the pointer and exceedancebox are removed.

(h) Engine Low Pressure Vibration Scale

The EICAS displays engine low pressure vibration as an analog display for theleft and right engines. Left engine low pressure vibration is shown on the left,and right engine low pressure vibration is shown on the right. The analogdisplay consists of a vertical scale, pointers that correspond to the engine lowpressure vibration values, and a white LP label centered below the scale.

The vertical scale is divided into five equal sections using four horizontal lines.The bottom of the scale represents 0 IPS and the top represents 2.5 IPS. Thevertical scale contains two bands as follows:

• The green band (normal) represents low pressure vibrations from 0 IPS to1.9 IPS

• The amber band represents low pressure vibrations from 1.9 IPS to 2.5IPS.



Two pointers are positioned on the vertical scale; one on the left side for theleft engine low pressure vibration, and the other on the right side for the rightengine low pressure vibration. Each pointer moves linearly between the topand bottom of the scale with respect to the color bands. When the engine lowpressure vibration exceeds the defined limits for a specific band, thecorresponding pointer changes color. If the engine low pressure vibrationexceeds the scale limits, the pointer parks at the respective end of the scale.For an invalid engine low pressure vibration, the pointer is removed.

(i) Engine High Pressure Vibration Scale

The EICAS displays engine high pressure vibration as an analog display forthe left and right engines. Left engine high pressure vibration is shown on theleft, and right engine high pressure vibration is shown on the right. Theanalog display consists of a vertical scale, pointers that correspond to theengine high pressure vibration values, and a white HP label centered belowthe scale.

The vertical scale is divided into five equal sections using four horizontal lines.The bottom of the scale represents 0 IPS and the top represents 2.5 IPS. Thevertical scale contains two bands as follows:

• The green band (normal) represents low pressure vibrations from 0 IPS to1.2 IPS

• The amber band represents low pressure vibrations from 1.2 IPS to 2.5IPS.

Two pointers are positioned on the vertical scale; one on the left side for theleft engine high pressure vibration, and the other on the right side for the rightengine high pressure vibration. Each pointer moves linearly between the topand bottom of the scale with respect to the color bands. When the enginehigh pressure vibration exceeds the defined limits for a specific band, thecorresponding pointer changes color. If the engine high pressure vibrationexceeds the scale limits, the pointer parks at the respective end of the scale.For an invalid engine high pressure vibration, the pointer is removed.

(j) Engine Annunciators

The EICAS displays annunciations for the engine modes, Automatic TakeoffThrust Control System, engine ignitors, FADEC in Control, and the thrustreversers. The engine annunciations are activated by signals coming from theFADEC via the DAU. These annunciations are displayed next to the N1 andITT indicators.

1 Engine Mode Annunciator

Engine mode annunciations are displayed at the top center of the engineinstrument section. The FADEC transmits the current mode for eachengine to the DAU via an ARINC 429 data bus. The DAU then sends thisinformation to the IAC for display processing.



For identical left and right engine modes, a single engine modeannunciation is display. If the left and right engine modes are different,both engine mode annunciations are displayed separately with exception ofthe ALT T/O-X annunciation. Dashes replace the annunciation when theleft and right alternated takeoff modes are different. Table 2-1-51 lists theengine mode annunciations which are displayed in cyan on the displayformat. For an invalid engine mode, six amber dashes (------) replace theannunciation.

Table 2-1-51. Engine Mode Annunciations

Mode EngineMode 3

EngineMode 2

EngineMode 1

EngineMode 0

Annunciation

Takeoff Mode 1 0 0 0 0 T/O-1



Alternate TakeoffMode 1

0 0 1 1 ALT T/O-1


0 1 0 0 ALT T/O-2


0 1 0 1 ALT T/O-3

Continuous 0 1 1 0 CON

Climb 0 1 1 1 CLB

Cruise 1 0 0 0 CRZ

Reversion Mode 1 0 0 1 R-MODE

PWT 1 1 1 1 PWT

2 Automatic Takeoff Thrust Control System Annunciator

The aircraft is equipped with an auxiliary power reserve that providesadditional thrust to the aircraft in case of an engine out. The EICASdisplays a green ATTCS annunciation on the display format when the leftand right Automatic Takeoff Thrust Control System (ATTCS) are armed.There is no invalid case for the ATTCS annunciation.

3 Ignition Annunciator

The aircraft is equipped with two ignitor plugs for each engine. Severalignition annunciations indicate the various states of the ignitors for eachengine. Table 2-1-52 lists the ignition annunciations which are displayed ingreen on the display format. There is no invalid case for the ignitionannunciation.



Table 2-1-52. Ignition Annunciation

Display FADEC AIgnition Enabled

FADEC AIgnition On

FADEC BIgnition Enabled

FADEC BIgnition On

IGNOFF

Not Enabled Off not enabled Off

IGNB

N/A Off N/A On

IGNA

N/A On N/A Off

IGNA B

N/A On N/A On

(Blank) Enabled Off Not Enabled Off

(Blank) Enabled Off Enabled Off

(Blank) Not Enabled Off Enabled Off

4 FADEC in Control Annunciator

The FADEC in Control annunciation indicates whether FADEC A or FADECB is in control of a particular engine. The IAC receives this data from theFADEC via the DAU. The FADEC in Control annunciation field displays agreen A if FADEC A is in control, or a green B if FADEC B is in control. Ifboth FADEC A and FADEC B indicate that they are in control, theannunciation field remains blank. The field also remains blank if bothFADEC A and FADEC B indicate that neither is in control. For an invalidindication, the FADEC in Control annunciation is removed.

5 Thrust Reverser Annunciator

The aircraft is equipped with a thrust reverser for each engine. The EICASdisplays a status indication of the thrust reverser for each engine. A greenREV annunciation is displayed to indicate that the thrust reverser is armedfor deployment when the thrust is boosted upon landing.

(2) Cabin and APU Section

The IAC receives various cabin information from the DAU which is then displayed inthe bottom left section of the display format. Cabin information includes cabinaltitude, cabin differential pressure, and cabin altitude rate.

The EICAS displays the current state of the APU, as well as, the current turbinespeed and temperature. The IAC receives this data from the DAU which is thendisplayed in the extreme lower left section of the display format, below the cabindata.



(a) Cabin Altitude Display

The cabin altitude is displayed in feet in a digital readout located at the top ofthe cabin and APU section. A white CAB ALT label annunciates the readout.The readout has a range from -1500 feet to +40000 feet with a resolution of100 feet. A white FT label is displayed next to the readout digits. Readoutdigits are displayed in the following colors:

• Green for altitudes up to 8100 feet

• Amber for altitudes from 8100 feet to 10,000 feet

• Red for 10,000 feet to 40,000 feet.

During normal conditions, the readout digits are not boxed. When the cabinaltitude exceeds a limit, the digits are boxed in the respective color of theexceedance. For an invalid cabin altitude, five amber dashes (-----) replacethe readout digits and the exceedance box is removed.

(b) Cabin Differential Pressure Display

The cabin differential pressure is displayed in pounds per square inch (PSI) ina digital readout, just below the cabin altitude. A white CAB P labelannunciates the readout. The readout range has a range of -0.5 PSI to 10.0PSI with a resolution of 0.1 PSI. A white PSI label is displayed next to thereadout digits. Table 2-1-53 lists the color and range for the cabin differentialpressure readout.

During normal conditions, the readout digits are not boxed. When the cabindifferential pressure exceeds a limit, the digits are boxed in the respectivecolor of the exceedance. For an invalid cabin differential pressure, threeamber dashes (--.-) replace the readout digits and the exceedance box isremoved.

Table 2-1-53. Cabin Differential PressureReadout

Range (PSI) Color

-0.5 to -0.3 Red

-0.3 to 0.0 Amber

0.0 to 8.0 Green (Normal)

8.0 to 8.4 Amber

8.4 to 10.0 Red



(c) Cabin Altitude Rate Display

The cabin altitude rate is displayed in feet per minute (FPM) in a digitalreadout, just below the cabin differential pressure. A white CAB RATE labelannunciates the readout. The readout range has a range of -2000 FPM to+2000 FPM with a resolution of 50 FPM. The readout digits are displayed ingreen. A white FPM label is displayed next to the readout digits. There is noexceedance limit defined for the cabin altitude rate. For an invalid cabinaltitude rate, five amber dashes (-----) replace the readout digits.

(d) APU Status Annunciator

The APU data is displayed on the bottom line in the cabin and APU section,below the cabin data. This data is displayed as either valid data, OFF, ordashes. If the APU is turned off, a green OFF annunciation is displayed andthe percent sign (%) and degree (°) symbols are removed.

(e) APU Turbine Speed Display

The APU turbine speed is displayed in revolutions per minute (RPM) in adigital readout. The readout has a range is of 0% RPM to 125% RPM with aresolution of one percent. A white % label is displayed next to the readoutdigits. Table 2-1-54 lists the color and range for the APU turbine speedreadout.

During normal conditions, the readout digits are not boxed. When the APUturbine speed exceeds a limit, the digits are boxed in the respective color ofthe exceedance. For an invalid APU turbine speed, three amber dashes (---)replace the readout digits and the exceedance box is removed.

Table 2-1-54. APU Turbine Speed

Range (% RPM) Color

0 to 95 Amber

96 to 104 Green (Normal)

105 to 110 Amber

111 to 125 Red



(f) APU Exhaust Temperature Display

The APU exhaust temperature is displayed in degrees Celsius in a digitalreadout. The readout has a range is of -73 °C to +977 °C with a resolution ofone degree. A white °C label is displayed next to the readout digits. Table2-1-55 lists the color and range for the APU exhaust temperature readout.

During normal conditions, the readout digits are not boxed. When the APUexhaust temperature exceeds a limit, the digits are boxed in the respectivecolor of the exceedance. For an invalid APU exhaust temperature, threeamber dashes (---) replace the readout digits and the exceedance box isremoved.

Table 2-1-55. APU Exhaust Temperature

ColorAPU Line

Contactor ClosedRange (°C)

APU Starter ContactorClosed (Start Mode)

Range (°C)

Green (normal) 0 to 680 0 to 838

Amber 681 to 717 839 to 884

Red 718 to 977 885 to 977

(3) Crew Alerting System Message Section

The IAC monitors the status of various aircraft and avionics systems on a continuousbasis and alerts the flight crew, as required, by displaying messages in the crewalerting system (CAS) message section of the display format. In addition toprioritizing and color coding the CAS messages, the IAC controls the messagetiming, flight crew acknowledgment, and message scrolling in order to declutter thedisplay.

Each IAC receives warning, caution, and advisory signals from the various aircraftand avionics systems and performs the same computations on the identical signalsindependently. Only messages from the IAC in control are displayed. The N2override switch on the maintenance allows CAS messages to be displayed while theaircraft is on the ground.

The CAS messages are monitored for validity. The controlling IAC performs achecksum on the transmitted CAS message and compares it to the checksum in thecross-side IAC. If a miscompare exists, a CAS message miscompare annunciation isdisplayed on the PFD. The controlling IAC continues to transmit messages to theEICAS display.



(a) CAS Messages

Sixteen lines of CAS messages can be displayed with 18 characters in eachline. A white box outlines the CAS message area on the left and bottomborders. The CAS messages are divided into three queues: warning, caution,and advisory. The CAS messages are displayed in each message queue inchronological order. New messages are inserted at the top of the respectivemessage queue. The most recent message appears at the top and the oldestmessage at the bottom of their respective message queue. The END statusmessage is always the last message displayed.

1 Warning Messages

Warning messages require immediate action from the flight crew. Warningmessages are displayed in red at the top of the display and are notscrolled off the screen. Unacknowledged warning messages flash at a rateof one second on, then 0.5 seconds off. A list of possible CAS warningmessages follows:

WARNING MESSAGES

APU FIRE AUTOPILOT FAIL

BAGG SMOKE BATT 1 OVTEMP

BATT 2 OVTEMP BLD 1 LEAK

BLD 2 LEAK BLD 1 OVTEMP

BLD 2 OVTEMP BLD APU LEAK

E1 ATTCS NO MRGN E1 LOW N1

E1 OIL LOW PRESS E2 ATTCS NO MRGN

E2 LOW N1 E2 OIL LOW PRESS

ELEC ESS XFR FAIL ENG 1 FIRE

ENG 2 FIRE ENG ATTCS FAIL

FUEL 1 TO LEVEL FUEL 2 LO LEVEL

GPWS ICE COND-A/I INOP

LAV SMOKE LG/LEVER DISAGREE

MAIN DOOR OPN NO TAKEOFF CONFIG

PIT TRIM 1 INOP PIT TRIM 2 INOP

SERVICE DOOR OPN SPS 1 INOP

SPS 2 INOP



2 Caution Messages

Caution messages indicate the possible need for action from the flightcrew. Caution messages are displayed in amber below the warningmessages. If there are no warning messages, caution messages arepositioned at the top of the display. If a warning message is generatedwhile caution messages are being displayed (acknowledged orunacknowledged), the warning message is inserted above the cautionmessages. Unacknowledged caution messages flash at a rate of onesecond on, then 0.5 seconds off.

A list of possible CAS caution messages follows:

CAUTION MESSAGES

115 VAC BUS OFF A/ICE SWITCH OFF

A/1 SW OFF ACCESS DOORS OPEN

AHRS 1 OVERHEAT AHRS 2 OVERHEAT

AIL SYS 1 INOP AIL SYS 2 INOP

AOA 1 HEAT INOP AOA 2 HEAT INOP

AP AIL MISTRIM AP ELEV MISTRIM

APU BLD VLV FAIL APU CNTOR CLSD

APU EXTBTL INOP APU FAIL

APU FIREDET FAIL APU FUEL LO PRESS

APU FUEL SOV INOP APU GEN OFF BUS

APU GEN OVLD APU OIL HI TEMP

APU OIL LO PRESS AURAL WARN FAIL

AUTO TRIM FAIL AUTOPILOT FAIL

BAGGAGE DOOR OPN BATT 1 OFF BUS

BATT 2 OFF BUS BKUP BATT OFF BUS

BLD 1 TEMP BLD 1 VLV FAIL

BLD 2 TEMP BLD 2 VLV FAIL

BRAKE DEGRADED BRAKE OVERHEAT

BRK INBD INOP BRK OUTBD INOP

CHECK PFD 1 CHECK PFD 2

CROSS BLD FAIL CROSS BLD SW OFF

DAU 1 ENG MISCMP DAU 2 ENG MISCOMP

DAU 1 SYS MISCMP DAU 1 WRN MISCMP

DAU 2 SYS MISCMP DAU 2 WRN MISCMP

DAU1 A FAIL DAU2 A FAIL



CAUTION MESSAGES

DC BUS 1 OFF DC BUS 2 OFF

DFDR FAIL E1 A/ICE FAIL

E1 ATS SOV OPN E1 CTL A FAIL

E1 CTL B FAIL E1 EXTBTL A INOP

E1 EXTBTL B INOP E1 FIREDET FAIL

E1 FUEL LO PRESS E1 FUEL LO TEMP

E1 FUEL SOV INOP E2 A/I FAIL

E2 ATS SOV OPN E2 CTL A FAIL

E2 CTL B FAIL E2 EXTBTL A INOP

E2 EXTBTL B INOP E2 FIREDET FAIL

E2 FUEL LO PRESS E2 FUEL LO TEMP

E2 FUEL SOV INOP ELEC EMERG ABNORM

ELEKBAY OVTEMP EMERG EXIT OPN

EMERG LT NOT ARMD EMRG BRK LO PRES

ENG NO TO DATA ENG REF A/I DISAG

ENG1 REV DISAGREE ENG1 REV FAIL

ENG1 TLA FAIL ENG2 REV DISAGREE

ENG2 REV FAIL ENG2 TLA FAIL

ESS BUS 1 OFF ESS BUS 2 OFF

FLAP FAIL FUEL IMBALANCE

FUEL TANK LO TEMP FUEL XFEED FAIL

FUELING DOOR OPN GEN 1 OFF BUS

GEN 1 OVLD GEN 2 OFF BUS



GEN 4 OVLD GPWS INOP

HS VLV 1 FAIL HS VLV 2 FAIL

HYD SYS 1 FAIL HYD SYS 1 OVHT

HYD SYS 2 FAIL HYD SYS 2 OVHT

IC 1 OVERHEAT IC 1 WOW INOP

IC 2 OVERHEAT IC 2 WOW INOP

IC BUS FAIL ICE DET1 FAIL



CAUTION MESSAGES

ICE DET2 FAIL ICE DETECTORS FAIL

LG AIR/GND FAIL NO ICE A/ICE ON

OVER WG ICE DET OVER WG ICE INOP

OXYGEN LO PRESS PACK 1 OVHT

PACK 1 OVLD PACK 1 VLV FAIL

PACK 2 VLV FAIL PACK 2 OVHT

PACK 2 OVLD PITOT 1 INOP

PITOT 2 INOP PITOT 3 INOP

PRESN AUTO FAIL RAM AIR VLV FAIL

RUD HDOV PROT FAIL RUDDER OVERBOOST

RUDDER SYS1 INOP RUDDER SYS2 INOP

SHED BUS 1 OFF SHED BUS 2 OFF

SPBK LVR DISAGREE SPOILER FAIL

SPS ADVANCED STAB A/ICE FAIL

STEER INOP STICK PUSHER FAIL

TAT 1 HEAT INOP TAT 2 HEAT INOP

W/S 1 HEAT FAIL W/S 2 HEAT FAIL

WG1 A/ICE FAIL WG2 A/ICE FAIL

WG A/ICE ASYMMETRY WINDSHEAR INOP

YAW DAMPER FAIL

3 Advisory Messages

Advisory messages indicate future attention of the flight crew. Advisorymessages are displayed in cyan in the space remaining, after the warningand caution messages. If a warning or caution message is generatedwhile advisory messages are being displayed (acknowledged orunacknowledged), the warning and/or caution message is inserted abovethe advisory messages. Advisory messages are automaticallyacknowledged after five seconds. Unacknowledged advisory messagesflash at a rate of one second on, then 0.5 seconds off.



A list of possible CAS advisory messages follows:

ADVISORY MESSAGES

AHRS1 BASIC MODE AHRS2 BASIC MODE

APU FUEL SOV CLSD BLD 1 VLV CLSD

BLD 2 VLV CLSD BRAKE DEGRADED

CHECKLIST MISMATCH CMC FAIL

CROSS BLD OPEN DAU1 B FAIL

DAU2 B FAIL DAU 1 REVERSION

DAU 2 REVERSION DU 1 FAN FAIL

DU 2 FAN FAIL DU 3 FAN FAIL

DU 4 FAN FAIL DU 5 FAN FAIL

E1 ADC DATA FAIL E1 FADEC FAULT

E1 FUEL IMP BYP E1 FUEL SOV CLSD

E1 HYD PUMP FAIL E1 HYD SOV CLSD

E1 IDL STP FAIL E1 OIL IMP BYP

E2 ADC DATA FAIL E2 FADEC FAULT

E2 FUEL IMP BYP E2 FUEL SOV CLSD

E2 HYD PUMP FAIL E2 HYD SOV CLSD

E2 IDL STP FAIL E2 OIL IMP BYP

FLAP LOW SPEED FUEL XFEED OPEN

GEN1 BRG FAIL GEN2 BRG FAIL

GEN3 BRG FAIL GEN4 BRG FAIL

HYD PUMP SELEC OFF HYD1 LO QTY

HYD2 LO QTY ICE CONDITION

IC 1 FAN FAIL IC 2 FAN FAIL

PACK 1 VLV CLSD PACK 2 VLV CLSD

4 Status Messages

The only status message used on this aircraft is the END message whichis displayed in white. This message is displayed at the end of the lastmessage queue. If no CAS messages are active, the END message isdisplayed at the top of the message queue and can not be scrolled off thetop of the display. If 15 CAS messages are active and none of themessages are scrolled off the top of the display, the END message isdisplayed on line 16 in place of the status line message. This is the onlyexception to the reserved status for line 16.



(b) Message Acknowledge

When new warning and caution messages are enabled, their status is set tounacknowledged. Unacknowledged warning and caution messages flash at arate of one second on, then 0.5 seconds off when displayed. Warning andcaution messages are acknowledged by pushing the applicable masterwarning and master caution annunciator/switch located in the glare shield.When a warning and/or caution message is acknowledged, the accompanyingaural tone is turned off.

All advisory messages automatically become acknowledged approximately fiveseconds after the messages first appear on the display format. Theacknowledged/unacknowledged status of any message not being displayedremains unchanged.

(c) Message Status Line

The CAS display reserves one line (line 16) as a message status line. Thepurpose of the status line is to indicate the existence of undisplayed caution oradvisory messages and their location relative to the currently displayedmessages. The status line consists of a parameter on each side of theMESSAGE label which indicates the total number of undisplayed messages.The left parameter indicates messages scrolled off the top of the display, andthe right parameter indicates messages scrolled off the bottom of the display.Arrows indicate the direction the messages are scrolled. The END message isnot included in the count of messages scrolled off the bottom of the display.

The status line remains blank if there are no undisplayed messages. If the 15available message positions are full such that a new message occurs but isundisplayed, the entire status line flashes until that message is displayed. Theparameters, arrows, and MESSAGE label are displayed in amber.

(d) Message Scrolling

The flight crew can scroll up and down through active caution and advisorymessages which do not appear on the display. Warning messages are notscrolled. A control knob located on the EICAS display bezel is used forscrolling through the messages. The control knobs provides the followingactions:

• Each clockwise click of the knob scrolls the caution/advisory messagequeue up one line

• Each counter clockwise click of the knob scrolls the caution/advisorymessage queue down one line

From their normal display position, messages can only be scrolled up.Scrolling up causes the displayed messages to move in relation to their normaldisplay position, thus removing the most recent message in the messagequeue. Scrolling up is inhibited if the message at the message queue isunacknowledged or the END message is displayed. The message queue canbe cleared by scrolling up until the END message is at the top of the queue.



Scrolling down provides a means of returning scrolled up messages to theirnormal display position. Scrolling down is inhibited if the message at thebottom of the CAS display is unacknowledged. The message queue canscrolled down until the most recent message is at the top of the queue.

If acknowledged caution/advisory messages are scrolled off the display and anew message is activated, the respective message queue is automaticallybrought into view. The new message is now displayed at the top of themessage queue.

(e) Message Inhibit

Certain flight phases require a message inhibit function because the flight crewcan not be distracted by enabling or disabling CAS messages. The logic thatdetermines which messages are inhibited is integrated into the messagedisplay logic. If a message is inhibited for an active inhibit phase, the IACquits processing its display logic, thus inhibiting the associated CAS messageon the display. When the message inhibit function is disabled, the IAC returnsto normal message processing to display any enabled CAS message, orremove any disabled CAS message. Table 2-1-56 lists the parameters thatinhibit the CAS messages.

NOTE: The N2 override switch on the maintenance panelallows the flight crew to override the N2 inhibitmessages. This allows the flight crew to view theinhibited messages that will be displayed when N2reaches a predetermined value.

Transitioning into a message inhibit phase has no effect on the currentmessage status. If a message is unacknowledged during the transition, themessage continues to flash until it is acknowledged. A CAS message isinvalid if it can not be acknowledged. For an invalid CAS, all messages areremoved and a red X is displayed on the CAS message section.

Table 2-1-56. Message Inhibit Parameters

TAKEOFF PARAMETER SET VALUE RESET VALUE

CAS V1 to 15 KTS < 60 KTS

WOW ON GND

RADIOALTITUDE

> 400 FT

APPROACH RADIOALTITUDE

TRANSITIONFROM >200 FT

TO <200 FT

WOW ON GND



(4) Flight Control Information Section

The EICAS displays information on landing gear status, flap position, and groundspoiler position in the upper portion of the flight control information section. TheEICAS displays pitch, roll, and yaw trim information in the lower portion of the flightcontrol information section.

(a) Landing Gear Display

The DAU sends a series of discretes to the IAC indicating the current state ofthe landing gear proximity sensors and landing gear lever. Logic is applied tothese discrete signals, and the status of each landing gear is displayed in abox. These boxes indicate from left to right the status of the left, the nose,and the right landing gear, respectively. A white LDG GEAR legend appearsabove the readouts.

Table 2-1-57 specifies the color of the box and annunciation for each gear andlever position listed in the table. If the aircraft is not properly configured fortakeoff, the landing gear annunciation and box are displayed in red. For aninvalid landing gear position, an amber dash (-) is displayed in thecorresponding readout and the outline box is displayed in amber.

Table 2-1-57. Landing Gear Positions

Landing GearControl Lever

Position

LandingGear

Position

StatusAnnunciation

Box/AnnunciationColor

Up Up UP White

Down Down DN Green

Down Up DN Red (Warning)

Up Down UP Red (Warning)

Up or Down In Transition Hash Marks Amber

(b) Flap Position Display

The DAU provides the IAC with the cardinal flap positions via an ARINC 429data bus. The EICAS uses this data to display a digital readout of theindicated flap position. The flap position is displayed in units of degrees ofsurface position.

The digital readout is outlined by a box, and is located below the landing gearreadouts. A white FLAPS legend is displayed next to the readout. Thereadout displays cardinal flap positions of 0, 9, 22, 33, and 45 degrees. Adash is displayed for flap positions in between the cardinal positions. Table2-1-58 specifies the color of the box and readout digits for each flap position.If the aircraft is not properly configured for takeoff, the flaps position readoutand box are displayed in red.



If the left flap position is invalid, then the right flap position is latched as theonly data source and remains latched until the IAC is re-powered. If both flappositions are invalid, the flap position is displayed as invalid. For an invalidflap position, an amber dash (-) replaces the readout digits and the outline boxis displayed in amber. If either flap control unit is invalid, the readout displaysa flap position in the range of 0° to 47°.

Table 2-1-58. Flap Positions

FlapPosition

DisplayedFlap Position

Box Color Digits Color

Invalid - Amber Amber

0 0 White (See Note) White (See Note)

1°-8° - White Green

9° 9 White (See Note) White (See Note)


22° 22 White (See Note) Green (See Note)


33° 33 White Green


45° 45 White Green

NOTE: The 0, 9,° and 22° cardinal positions are displayed in redinside a red outline box when

the aircraft is on the ground (weight on wheels).

(c) Ground Spoiler Display

The DAU sends the ground spoiler position to the IAC to indicate the currentstate of the ground spoilers. The EICAS then displays the ground spoilerstatus as either open or closed in a digital readout, located just below the flapsposition readout. A white SPLRS legend is displayed next to the readout.

A white box outlines the readout digits for normal conditions. If the left andright inboard and outboard switches all indicate a closed status, the readoutdisplays a white CLD annunciation. If any of the four inboard and outboardswitches indicate an open status, the readout displays a green OPNannunciation. When any or all of the four switches indicate an open or closedstatus with the aircraft on the ground (weight-on-wheels) and a thrust leverageangle of less than 60 degrees, the readout displays the corresponding CLD orOPN annunciation in red inside a red outline box. If the aircraft is not properlyconfigured for takeoff, the ground spoiler annunciation and box are displayedin red. There is no invalid display for ground spoiler data.



(d) Pitch Trim Display

The IAC receives pitch trim information from the aircraft in an analog format.The EICAS then displays a vertical scale with pointer and a digital readout toindicate the pitch trim position. A white PITCH label annunciates the pitch trimdisplay. Left pitch trim sensors provide the primary data source for pitch triminformation. Right pitch trim sensors provide a secondary data source forpitch trim information.

The white vertical scale is divided into four equal segments, with long dasheson the ends and in the middle and short dashes located at the intermediatescale positions. The vertical scale is scaled linearly with the bottomrepresenting 100% pitch trim down (4°), and the top representing 100% pitchtrim up (-10°). The pitch angle is measured on the leading edge of thestabilizer. A pointer moves vertically along the scale to indicate the pitchangle. A green band is positioned above the center of the vertical scale from -4° to -8° to represent the take-off pitch angle.

A digital readout of the pointer position is displayed next to the vertical scale.The readout has a range of 4° (100% pitch down) to -10° (100% pitch up) witha resolution of one degree. A box outlines the readout digits. If the left pitchtrim position is greater than or equal to 1°, then DN is annunciated above thereadout. For left pitch angles less than or equal to -1°, the minus sign (-) isdisplayed and UP is annunciated above the readout. If the pitch angle isbetween -1° and +1°, both the readout and UP/DN annunciations remainblank.

Table 2-1-59 specifies the color of the pointer, readout digits and box, and theUP/DN annunciations when weight on wheels indicates an on ground status.The readout digits, UP/DN annunciations, and pointer are displayed in greenand the box in white for an in air status. If the aircraft is not properlyconfigured for takeoff, the pointer is displayed in red. For an invalid pitch trimposition, the pointer is removed from the display and three amber dashes (---)replace the readout digits.

Table 2-1-59. Pitch Trim Display

PitchTrim

Position

EngineThrust

Lever Angle

EngineThrust LeverAngle Valid

Weight OnWheels

Valid

PointerColor

UP/DNAnnunciation

Box/DigitsColor

-4°-8° ≥60° Valid On Ground Red Red Red

-4°-8° <60° Valid On Ground Amber Amber Amber



(e) Roll Trim Display

The IAC receives roll (aileron) trim information from the aircraft in an analogformat. The EICAS then displays an arc with pointer to indicate the ailerontrim position. The aileron trim arc is displayed to the left of the pitch trimvertical scale. A white ROLL label annunciates the roll trim display.

The white aileron trim arc is divided into four equal arc segments, with longdashes on the ends and in the middle and short dashes located at theintermediate scale positions. The arc is scaled linearly with the left edgerepresenting -100% left aileron trim (left wing down), and the right edgerepresenting +100% right aileron trim (right wing down).

A green pointer sweeps the scale on the inside of the arc to indicate theaileron trim position. If aileron trim exceeds the limits of the scale, the pointerparks itself at the respective end of the scale. An amber mistrim arrowindicates when the aircraft is not trimmed correctly. The mistrim arrow pointsin the direction needed to correct the trim and moves with the pointer. For aninvalid aileron trim position, the pointer and mistrim arrow are removed fromthe display.

(f) Yaw Trim Display

The DAU transmits the current yaw (rudder) trim position to the IAC via anARINC 429 data bus. The EICAS then displays a horizontal scale with pointerto indicate the yaw trim position. The yaw trim scale is positioned below theroll trim arc. A white YAW label annunciates the yaw trim display.

The horizontal scale is divided into four equal segments, with long dashes onthe ends and in the middle and short dashes located at the intermediate scalepositions. The horizontal scale is scaled linearly with the left edgerepresenting -100% left rudder trim, and the right edge representing +100%right rudder trim up. A green pointer moves horizontally along the scale toindicate the rudder trim tab position. If the rudder trim exceeds the limits ofthe scale, the pointer parks itself at the respective end of the scale. For aninvalid rudder trim position, the pointer is removed from the display.



M. EICAS Test Mode

The system test format provides for a failure mode annunciation and familiarization of theEICAS display format. The test mode is functional while weight-on-wheels is sensed andthe airspeed is less than 50 knots. Pushing and holding the TEST button on theappropriate DC-550 Display Controller initiates the test mode for the corresponding MFD.The test page remains displayed as long as the test mode is active. Figure 2-1-63 showsthe failure mode annunciation and familiarization format for the EICAS display. Displayformats not specifically mentioned do not change.

The current engine mode is forced into a split display. The CAS display is forced into atest format (a red X is drawn through the CAS message area). If the English/Metric strapindicates metric, then the fuel flow and fuel tank quantity labels are displayed as metricunits.

The following display items are displayed in their invalid state:

N1 ITT

N2 Fuel Flow

Fuel Quantity Oil Pressure

Oil Temperature LP Variation

HP Variation Cabin Altitude

Cabin Differential Pressure Cabin Rate

APU Status Landing Gear

Flaps Pitch Trim

Roll Trim Yaw Trim

Pitch Trim



AD-51248@

---.-

LDG GEAR

FLAPS

SPLRS CLD

APU

CAB ALT

CAB P

CAB RATE

C

FT

PSI

FPM

---

-----

-----

%

VIB

LP HP

-- --

TEMPPRESS

---

N1

ITT---- ----

---.-

N2---.- %%

FF---- PPHPPH

FQ LBLB

--

ROLL

YAW

PITCH

INVALID N1 AND ITTDIGITAL READOUTS

(AMBER DASHES)

INVALID EICASPARAMETERS(AMBER DASHES)

INVALID OIL PRESSUREAND TEMPERATURE

(AMBER DASHES)

----

---.-

----

----

---.- ------- ------ ---.-

IGN

OFF

IGN

OFF

OIL

---

--.-

---

INVALID LANDINGGEAR AND FLAPS(AMBER DASHES)

INVALID CASMESSAGE FIELD

INVALID ENGINE MODESAND DIGITAL READOUTS

(AMBER DASHES)

INVALID N2, FUEL FLOW AND FUEL QTY

(AMBER DASHES)

INVALID APU TURBINE SPEEDAND EXHAUST TEMPERATURE

(AMBER DASHES)

INVALID CABINDATA

(AMBER DASHES) --

NOTES: 1. THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS.

2. POINTERS ARE REMOVED FROM INVALID ANALOG SCALES.

---.- ------- ------ ---.-

Figure 2-1-63. EICAS Familiarization Display



N. Reversionary Modes

(1) Symbol Generator Reversion

The EDS uses two symbol generators, one in each IAC. Each symbol generator iscapable of driving three separate display formats (PFD, MFD, and EICAS). Thesystem is configured such that under normal operating conditions the pilot’s symbolgenerator (IAC No. 1) provides the PFD format on display unit (DU) No. 1, the MFDformat on DU No. 2, and the EICAS format on DU No. 3. The copilot’s symbolgenerator (IAC No. 2) provides an MFD format on DU No. 4 and a PFD format onDU NO. 5. Figure 2-1-63 shows the interface diagram for the symbol generatorreversion mode.

Each IAC uses its own-side sensor information for its respective PFD and MFDdisplays. If symbol generator No. 1 fails (PFD No. 1, MFD No. 1, and the EICASdisplay are blank), pushing the SG reversion switch on the pilot’s reversionary panelallows symbol generator No. 2 to drive all five display units. The outer display units(DU No. 1 and DU No. 5) display identical PFD formats. Displays units No. 2 andNo. 4 display identical MFD formats. The center display unit displays the EICASformat. Likewise, if symbol generator No. 2 fails, pushing the SG switch on thecopilot’s reversionary panel allows symbol generator No. 1 to drive all five displayunits in the same manner.

(2) Sensor Reversion

If an AHRS or MADC fails, the system can be reconfigured to display the remaininggood sensor on both sides of the cockpit. The remaining good sensor is selectedusing the AHRS or ADC switches on the reversionary panel.

Each DAU has two data channels: channel A and channel B. Each DAU providesboth channels of data to each IC-600 IAC. Normally, the EICAS uses the channel Adata from its own-side DAU and channel A data from the cross-side DAU. If channelA of DAU No. 1 fails, the DAU 1 reversion switch on the EICAS panel is toggled tothe REV position, resulting in both symbol generators looking at channel B data fromDAU No. 1. Likewise, if channel A of DAU No. 2 fails, the DAU 2 reversion switch istoggled to the REV position, resulting in both symbol generators looking at channel Bdata from DAU No. 2. The relaxed state of the DAU reversion switch is the normalstate. Pushing the switch to its inward position constitutes the reversion state.

Reversion control is accomplished as follows:

(a) ADC Reversion

MADC No. 1 is the pilot on-side/co-pilot cross-side MADC. MADC No. 2 is thepilot cross-side/copilot on-side MADC. If MADC reversion indicates normal,the IC-600 IACs uses the on-side MADC data for displaying air datainformation. If MADC reversion indicates reversion for the pilot’s side, thepilot’s IAC uses cross-side MADC data for displaying air data information. IfMADC reversion indicates reversion for the copilot’s side, the copilot’s IACuses cross-side MADC data for displaying air data information. Figure 2-1-64shows the interface diagram for the ADC reversion mode.



(b) AHRS Reversion

AHRS No. 1 is the pilot on-side/co-pilot cross-side AHRS. AHRS No. 2 is thepilot cross-side/copilot on-side AHRS. If AHRS reversion indicates normal, theIC-600 IACs uses the on-side AHRS data for displaying air data information. IfAHRS reversion indicates reversion for the pilot’s side, the pilot’s IAC usescross-side AHRS data for displaying air data information. If AHRS reversionindicates reversion for the copilot’s side, the copilot’s IAC uses cross-sideAHRS data for displaying air data information. Figure 2-1-65 shows theinterface diagram for the AHRS reversion mode.

(c) DAU Reversion

Each DAU has two data channels that provide engine data to each IAC. TheIACs use engine data to display left and right primary engine information onthe EICAS display. The DAUs operate in either the normal mode or thereversion mode to provide the engine data. Figure 2-1-66 shows the interfacediagram for the DAU reversion mode. Table 2-1-60 specifies which channel ofeach DAU is the primary source of engine data in both the normal mode andthe reversion mode:

Table 2-1-60. DAU Reversion States

EICAS ReversionSwitch

IC-600 IAC

Pilot Copilot

DAU 1Reversion

Select

DAU 2Reversion

Select

LeftEngine

Data

RightEngine

Data

LeftEngine

Data

RightEngine

Data

Normal Normal DAU 1Channel A

DAU 2Channel A

DAU 1Channel A

DAU 2Channel A

Reversion Normal DAU 1Channel B

DAU 2Channel A

DAU 1Channel B

DAU 2Channel A

Normal Reversion DAU 1Channel A

DAU 2Channel B

DAU 1Channel A

DAU 2Channel B

Reversion Reversion DAU 1Channel B

DAU 2Channel B

DAU 1Channel B

DAU 2Channel B

DAU 1 = Pilot’s DAU, DAU 2 = Copilot’s DAU



(3) Display Unit Reversion

The MFD reversion switches on the reversionary panels control the display unitreversionary operation. The reversion switches allow the PFD or EICAS format to bemoved from a failed DU (blank or flashing) to good DU. The lost display format froma faulty DU is moved to either DU No. 2 or DU No. 4 by selecting the desired formatwith the MFD reversion switch. The system gives the PFD top priority followed bythe EICAS display, with the least priority given to the MFD. Each IAC then drivesthe display formats listed in Table 2-1-61 depending on the positions of the MFD andSG reversion switches on the reversionary panels.

Table 2-1-61. DU Reversion States

Reversionary PanelSwitch Positions Display Unit Formats

IAC MFD1 MFD2DU No.

1DU No.

2DU No.

3DU No.

4DU No.

5

NORM NORM NORM PFD1 MFD1 EICAS1 MFD2 PFD2

NORM NORM PFD PFD1 MFD1 EICAS1 PFD2 ---

NORM NORM EICAS PFD1 MFD1 --- EICAS2 PFD2

NORM PFD NORM --- PFD1 EICAS1 MFD2 PFD2

NORM PFD PFD --- PFD1 EICAS1 PFD2 ---

NORM PFD EICAS --- PFD1 --- EICAS2 PFD2

NORM EICAS NORM PFD1 EICAS1 --- MFD2 PFD2

NORM EICAS PFD PFD1 EICAS1 --- PFD2 ---

NORM EICAS EICAS PFD1 EICAS1 --- EICAS2 PFD2

SG1 NORM NORM PFD1 MFD1 EICAS1 MFD1 PFD1

SG1 NORM PFD PFD1 MFD1 EICAS1 PFD1 ---

SG1 NORM EICAS PFD1 MFD1 --- EICAS1 PFD1

SG1 PFD NORM --- PFD1 EICAS1 MFD1 PFD1



Table 2-1-61. DU Reversion States

SG1 PFD PFD --- PFD1 EICAS1 PFD1 ---

SG1 PFD EICAS --- PFD --- EICAS1 PFD1

SG1 EICAS NORM PFD1 EICAS1 --- MFD PFD1

SG1 EICAS PFD PFD1 EICAS1 --- PFD1 ---

SG1 EICAS EICAS PFD1 EICAS1 --- EICAS1 PFD1

SG2 NORM NORM PFD2 MFD2 EICAS2 MFD2 PFD2

SG2 NORM PFD PFD2 MFD2 EICAS2 PFD2 ---

SG2 NORM EICAS PFD2 MFD2 --- EICAS2 PFD2

SG2 PFD NORM --- PFD2 EICAS2 MFD2 PFD2

SG2 PFD PFD --- PFD2 EICAS2 PFD2 ---

SG2 PFD EICAS --- PFD2 --- EICAS2 PFD2

SG2 EICAS NORM PFD2 EICAS2 --- MFD2 PFD2

SG2 EICAS PFD PFD2 EICAS2 --- PFD2 ---

SG2 EICAS EICAS PFD2 EICAS2 --- EICAS2 PFD2

NORM indicates no reversion is selected--- indicates the DU is off, DU reversion is selectedSG1 indicates SG No. 1 reversion is selected, SG No. 2 is driving all

active DUsSG2 indicates SG No. 2 reversion is selected, SG No. 1 is driving all

active DUsPFD indicates PFD reversionEICAS indicates EICAS reversion



16

DU NO. 1 - PFD 1

22

130J1

878889

909192

DU POWER DN

PORT SEL 3PORT SEL 2PORT SEL 1

ID NO. 3ID NO. 2ID NO. 1

IC-600 NO. 2C190J2B

48 SG REV

DC-550 NO. 2 C115J1

777872

CS MFD=EICAS

COPILOTS IC-600REVERSIONSWITCH

PILOTS IC-600REVERSIONSWITCH ANNUNCIATOR

POWER

LIGHTINGGND

LIGHTINGGND

79CS MFD=PFD

OS MFD=EICASOS MFD=PFD

IC-600 NO. 1190J2B

48 SG REV

DC-550 NO. 1 115J1

777872

CS MFD=EICAS 79CS MFD=PFD

OS MFD=EICASOS MFD=PFD

PFD

NRM

EICAS

MFD1REVSW

DU NO. 2 - MFD 1

22

131J1

878889

909192

DU POWER DN



DU NO. 3 - EICAS

22

132J1

878889

909192

DU POWER DN



DU NO. 4 - MFD 2

22

C131J1

878889

909192

DU POWER DN



PFDNRM

EICAS

MFD2REVSW

DU NO. 5 - PFD 2

22

C130J1

878889

909192

DU POWER DN



AD-39748-R2@

28 V DCBUS 2

28 V DCBUS 1

K3

K1

K2

15

35

34 DC/ICBUS OUT

DC/ICBUS IN

15

3534

190J2A

H

L

H

L

115J1

C115J1

DC/ICBUS OUT

H

L

H

LDC/IC

BUS IN

C190J2A

63

LATCHINGSWITCHES

16

CROSS-SIDE SG REV

Figure 2-1-64. Symbol Generator Reversion Mode Interface Diagram



DC-550 NO. 1

MADC REV

115J1

74NORMREV

PILOT'SMADC REVSWITCH

AZ-850 MADC NO. 1

IC-600 IAC NO. 1

429 CH10 IN

190J2A

43

15

44

16

429CH 3IN

190J2B

IC-600 IAC NO. 2429

CH 3 IN

C190J2B

23

43

24

44429CH 10 IN

C190J2A

429 BUSNO. 1 OUT

9J1

6061

AH-800 NO. 1

K4

K5429 CH 1 IN

1J1B

AH-800 NO. 2

429 CH 1 IN

C1J1B

K4K5

OPTIONAL P-870WX RADAR

59J1

RCZ-851E NO. 1

429 CH 2 IN

143J1

10080

6999

143J1

429 CH 1 IN86

429 BUSNO. 3 OUT

9J1

6667

RCZ-851E NO. 2

429 CH 2 IN

C143J1

1008069

99

C143J1

429 CH 1 IN86

TO HORIZONTALSTABILIZER

CONTROL UNIT 9J1

6364

TO FADEC 1A

TO FADEC 1A

RESERVED

TO FADEC 2A

TO FADEC 2A

429 BUSNO. 2 OUT

TO PRESSUREDIGITAL CONTROLLER 9J1

6869

TO STALL PROTECTIONCOMPUTER

TO GROUND PROXWARNING SYSTEM

429 BUSNO. 4 OUT

DC-550 NO. 2

MADC REV

C115J1

74

AZ-850 MADC NO. 2

ARINC 429 BUSNO. 1 OUT

C9J1

6061


C9J1

6667

TO HORIZONTALSTABILIZERCONTROL UNITC9J1

6364

TO FADEC 1BTO FADEC 1B

RESERVED

TO FADEC 2BTO FADEC 2B


TO PRESSUREDIGITAL CONTROLLERC9J1

6869

TO STALL PROTECTIONCH3 COMPUTER

TO WINDSHEARCOMPUTER

TO OPTIONALFMS


AD-50640-R1@

TO OPTIONALFMS/GPS

TO OPTIONALGPS

tr

3536

AP 429CH 1 IN

COPILOT'SMADC REV

SWITCH

LATCHING SWITCHES

3435

190J2A

2324

DC/IC BUSIN

DC/IC BUSOUT

1516

C190J2A

3435

DC/IC BUSOUT

H

L

H

L

HL

H

L

H

L

H

L

H

L

H

L

H

L

H

LH

L

H

L

H

L

H

L

H

L

H

L

H

L

DC/ICBUS IN

HL

HL

H

L

H

L

H

L

H

L

Figure 2-1-65. MADC Reversion Interface Diagram



429 CH 4 IN2930

1516

429 CH 11 IN8384

PILOT'S IC-600

429 CH 4 IN2930

1516

429 CH 11 IN8384

COPILOT'S IC-600C190J2A190J2A

G8G7

E5E6

COPILOT'S AH-800

429 OUT 1

429 OUT 2

C1J1B

G8G7

E5E6

PILOT'S AH-800

429 OUT 1

429 OUT 2

1J1B

73

3435

COPILOT'S DC-550

AHRS 1 SEL

DC/IC BUSOUT

C115J1

73

3435

PILOT'S DC-550

AHRS 2 SEL

DC/IC BUSOUT

115J1

AD-50095@

K12K13

F14F15

429 OUT 3

429 OUT 4

K12K13

F14F15

429 OUT 3

429 OUT 4 SPS,GPWSFMS

SPS,GPWSFMS

59J1-K59J1-L

WEATHERRADAR

4342

DAU2 DA-800

429 IN 2

137J1B

AHRSREVERSION

SWITCH

HL

AP 429 CH 0 IN

DC/IC BUSIN H

L

3334

DC/IC BUSINH

L

Figure 2-1-66. AHRS Reversion Mode Interface Diagram



429 CH 9 IN81

82

429 CH 8 IN4142

PILOT'S IC-600

190J2A

5150

PILOT'S DAU CH A

429 CH2 OUT

136J1B

76

75

PILOT'S DC-550

DAU B SEL

115J1

PILOTSDAU

REVERSIONSWITCH

AD-50639@

DAU A SEL3534

429 CH 7 IN3132

429 CH 6 IN3132

16

15

190J2A

115J1

5150

PILOT'S DAU CH B

429 CH2 OUT

136J2B

429 CH 9 IN81

82

429 CH 8 IN4142

COPILOT'S IC-600C190J2A

51

50

COPILOT'S DAU CH A

429 INPUT

429 CH2 OUT

137J1B

76

75

COPILOT'S DC-550

DAU B SEL

C115J1

DAU A SEL 3534

429 CH 7 IN3132

429 CH 6 IN3132

C190J2B

16

15

C190J2A

C115J1

51

50

COPILOT'S DAU CH B

429 CH2 OUT

137J2B

3738

429 OUTPUT

43

42

4041

429 CH1 IN

429 CH2 OUT

C190J2A

4041

137J1B

TO OTHERAIRCRAFTSYSTEMS(SEE SECTION 3)

429 CH1 OUT

NORM

REV

COPILOTSDAU

REVERSIONSWITCH

LATCHING SWITCHES

DC/IC BUS

L

H DC/IC BUSIN

190J2A

190J2A

190J2A

136J1B

DC/IC BUS

38

37

4041

L

HDC/IC BUSIN

TO OTHERAIRCRAFTSYSTEMS

(SEE SECTION 3)

Figure 2-1-67. DAU Reversion Mode Interface Diagram





2.2 AHZ-800 ATTITUDE HEADING REFERENCE SYSTEM . . . . . . . . . . . . . . . . . . . . . . . 2-2-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-1


A. AH-800 Attitude Heading Reference Unit (AHRU) . . . . . . . . . . . . . . . 2-2-2

B. Memory Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-4

C. FX-600 Flux Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-5

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-6

A. Pilot’s AHRS Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-6

B. Copilot’s AHRS Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-6

C. Modes of Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-10

(1) Initialization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-10

(2) Full Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-10

(3) DG Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-10

(4) Basic Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-11

(5) Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-11

(6) Maintenance Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-12

D. Reversionary Switching . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-12

4. Fault Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-13


B. Power-On BITE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-14

C. Continuous BITE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-15

D. Fault Reaction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-16

(1) Critical . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-16

(2) Non-Critical . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-16

E. Maintenance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-16

F. Fault Storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-16

G. Flight Faults . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-16

(1) Fault Service Status . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-17

(2) Miscellaneous Status . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-17

(3) Fault Type . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-17

H. Ground Faults . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-17

I. Fault Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-18

(1) PFD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-18

(2) MFD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-20

(3) EICAS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-23




FIGURE/TITLE PAGE


Figure 2-2-1. AH-800 Attitude Heading Reference Unit . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-2

Figure 2-2-2. Memory Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-4

Figure 2-2-3. FX-600 Flux Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-5

Figure 2-2-4. Pilot’s AHRS Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-7

Figure 2-2-5. Copilot’s AHRS Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-8

Figure 2-2-6. PFD AHRS Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-19

Figure 2-2-7. MFD MAP MODE AHRS Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . 2-2-21

Figure 2-2-8. MFD PLAN MODE AHRS Failure Indications . . . . . . . . . . . . . . . . . . . . . . . 2-2-22

List of Tables

TABLE/TITLE PAGE

Table 2-2-1. AH-800 Attitude Heading Reference Unit Leading Particulars . . . . . . . . . . . . 2-2-3

Table 2-2-2. Memory Module Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-4

Table 2-2-3. FX-600 Flux Valve Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-5

Table 2-2-4. AH-800 ARINC 429 Output Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-9

Table 2-2-5. AH-800 Full Performance Accuracy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-11

Table 2-2-6. AHRS ARINC 429 Output Test Mode Data . . . . . . . . . . . . . . . . . . . . . . . . . . 2-2-12



SECTION 2.2AHZ-800 ATTITUDE HEADING REFERENCE SYSTEM

1. General

The PRIMUS® 1000 Avionics System installed in the Embraer 145 aircraft includes a newgeneration attitude heading reference system (AHRS) - the AHZ-800. The AHZ-800 design isbased on a fiber optic gyro developed by Honeywell for this class of system. The gyro isreferred to as an interferometric fiber optic gyro (IFOG). A dual AHZ-800 system is standard onthis aircraft.

The basic system is comprised of the following LRUs:

• AH-800 Attitude Heading Reference Unit (AHRU)

• Memory Module

• FX-600 Flux Valve.

In the Embraer 145 installation, a separate AHRU mode controller is not required. Thecontroller function is integrated within the cockpit using remote switches and a pushbutton.

The AHRU senses aircraft attitude and physical motion, and obtains long-term magnetic headinginformation from the flux valve. After performing the necessary computations, the AHRUcomputes the following outputs and transmits them on ARINC 429 buses.

• Attitude (pitch and roll)

• Magnetic Heading

• Angular Rate of Change (pitch, roll, and yaw)

• Linear Acceleration (longitudinal, lateral, and normal)

• Operating Mode Status.




A. AH-800 Attitude Heading Reference Unit (AHRU)

Figure 2-2-1 shows a graphical view of the AH-800 Attitude Heading Reference Unit(AHRU). The AHRUs are located in the avionics nose bay of the aircraft. Table 2-2-1provides items and specifications that are particular to the AHRU.

AD-32466@

Figure 2-2-1. AH-800 Attitude Heading Reference Unit



Table 2-2-1. AH-800 Attitude Heading Reference Unit Leading Particulars

Item Specification

Dimensions (maximum) . . . . . . . . 4 MCU per ARINC 600

• Height . . . . . . . . . . . . . . . . . 7.64 in. (194.06 mm)

• Width . . . . . . . . . . . . . . . . . 4.88 in. (123.95 mm)

• Length (from rear connector) . 15.12 in. (384.05 mm)

Weight . . . . . . . . . . . . . . . . . . . . 16.0 lb (7.32 kg)

Power Requirements . . . . . . . . . . 28 V dc, 45 W (max)


• Fan, FN-260 . . . . . . . . . . . . HPN 26006881-102

• Fan Filter . . . . . . . . . . . . . . HPN 26000790-101

• Memory Module, MM-260 . . . HPN 26006926-102

Mating Connectors:

• ARINC 600 Connector WithoutContacts (1) . . . . . . . . . . . . . HPN 10070295-102

• ARINC 600 20 AWG Contacts HPN 10072691-102




Mounting:

• Tray, MT-260, 4 MCU, withside mounted fan . . . . . . . . . HPN 26012346-101

The AHRU is the major component of the system. It contains the necessary powersupplies, sensors, and electronics to compute aircraft attitude, magnetic heading, rate ofchange, and acceleration forces. The AHRU outputs digital data for the electronic displaysystem (EDS), flight guidance system (FGS), flight management system (FMS), weatherradar antenna, and other aircraft systems as required. The sensors within the AHRUinclude fiber optic gyros, which sense angular motion around all three axes; andaccelerometers, which sense linear motion along all three axes. It is capable of360-degree displacement in the roll and heading axes, and ±85 degree displacement in thepitch axis.

The AHRU provides the excitation, current feedback control, and signal demodulationinterfaces for the flux valve.



B. Memory Module

Figure 2-2-2 shows a graphical view of the Memory Module and Table 2-2-2 provides itemsand specifications that are particular to the unit.

AD-32840@

Figure 2-2-2. Memory Module

Table 2-2-2. Memory Module Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . 0.72 in. (18.29 mm)

• Width . . . . . . . . . . . . . . . . . 1.16 in. (29.46 mm)

• Length (excluding cable) . . . . 2.16 in. (54.86 mm)

Weight (maximum) . . . . . . . . . . . 0.08 lb (0.04 kg)

Power Requirements . . . . . . . . . . Supplied by AHRU

User Replaceable Parts . . . . . . . . None

The memory module is located on the rear of the AHRU mounting tray and includes asystem EEPROM that is used to store such aircraft specific data as AHRU configurationdiscretes, mounting misalignment coefficients, and flux valve compensation coefficients.

Programming of the memory module is accomplished with a laptop computer. Memorymodule programming is only possible when the memory access mode select discrete isasserted, and the memory access interface is enabled (weight-on-wheels is at ground andTAS is less than 60 knots).

Memory module configuration data is programmed during aircraft installation of the AHRS.Subsequent replacement of either the memory module or the flux valve will require theEEPROM configuration data to be reprogrammed, and the flux valve to be recalibrated inaccordance with the procedures in MAINTENANCE PRACTICES, SECTION 4 of thismanual.



C. FX-600 Flux Valve

Figure 2-2-3 shows a graphical view of the FX-600 Flux Valves. The flux valves arelocated on the wing tips of the aircraft. Table 2-2-3 provides items and specifications thatare particular to the unit.

AD-32728@

Figure 2-2-3. FX-600 Flux Valve

Table 2-2-3. FX-600 Flux Valve Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . 2.06 in. (25.32 mm)

• Width . . . . . . . . . . . . . . . . . 3.60 in. (91.44 mm)

• Length . . . . . . . . . . . . . . . . 4.40 in. (111.76 mm)

• Bowl Diameter . . . . . . . . . . . 2.31 in. (84.07 mm)

Weight (maximum) . . . . . . . . . . . 1.44 lb (0.65 kg)

Power Requirements . . . . . . . . . . 28 V dc square wave

User Replaceable Parts . . . . . . . . Non-magnetic machine screws, No. 6-40 by 3/8-inchfillister head, HPN 1715115

Mating Terminals (6) . . . . . . . . . . HPN 0364-01

Mounting . . . . . . . . . . . . . . . . . . Non-magnetic machine screws, No. 6-40 by 3/8-inchfillister head, HPN 1715115

The FX-600 Flux Valve senses the horizontal portion of the earth’s magnetic field andprovides a long-term heading reference. The heading reference is processed by the AHRUto compute an inertially stabilized magnetic heading output. See Figure 2-? for locationwithin the aircraft.



3. Operation

A. Pilot’s AHRS Interface

See Figure 2-2-4.

The fundamental inputs to the AHRU are derived from the aircraft’s inertial motion. Thesystem contains three accelerometers, three fiber optic gyros, and a flux valve. Thesesensors are used to measure inertial motion of the aircraft to compute pitch and rollattitude, magnetic heading, angular rates, and linear acceleration. Additionally, the AHRUreceives true airspeed (TAS) from the MADC, via an ARINC 429 bus to compensate foracceleration induced errors.

As shown in Figure 2-2-4, the AHRU has the following:

• Memory access via the aircraft test connector.

• Mode control inputs from the following three cockpit mounted switches:

– SLAVED/DG toggle switch enables the operator to select between thesetwo modes. In the DG mode, the AHRU operates as a free gyro.

– When on the ground, the AHRS TEST momentary pushbutton switchcauses the AHRU to transmit the standard self-test output data.

– HEADING SLEW spring loaded, center off, toggle switch allows resettingof the heading when in the DG mode.

• Excitation to the FX-600 Flux Valve is required so it can sense the horizontal portion ofthe earth’s magnetic field to provide a long-term heading reference to the AHRU. Theexcitation is 28 V dc chopped to ground level at 400 Hz.

• All PRIMUS® 1000 LRUs that require AHRS data receive that data over the ARINC 429buses. For output data information, refer to Table 2-2-4.

• Connections to the MM-260 Memory Module to store such aircraft specific data asAHRU configuration discretes, mounting misalignment coefficients, and flux valvecompensation coefficients.

B. Copilot’s AHRS Interface

See Figure 2-2-5.

The copilot’s system is identical to the pilot’s system except that the ARINC 429 No. 3output is connected to the weather radar for antenna stabilization instead of the StallProtection Computer and GPWS.



H3H4H7H10

728

K12K13

AIRCRAFTTESTCONNECTOR

28 V DC FROM ESS DC BUS 1

DC GND28 V DC FROM BACKUP ESS DC BUS 1

HL

ARINC 429NO. 3 OUT

SLVD

DG

AHRSTEST

H6

A10

CW

CCW

H1

H4

H2

K15

H3

COM012

WOW GROUND = ON GROUND

H8H9

H10H12H13H11

HL

F6F7F8F9

F11

INHIBITPWR

RETURNCLOCK

DATA

MEM ACCESS DATA TXMEM ACCESS DATA RX

MEM ACCESS COMMMEM ACCESS MODE SEL

28 V DC PRIMARY28 V DC AUXILARY

28 V DC RETURN

DG MODE SELECT

TEST COMMAND

HEADING SLEW

ABCDEF

FLUX VALVE INPUT

FLUX VALVEEXCITATION

FLUX VALVE COMMON

TOMEMORY MODULE

MEMORYMODULE

A B C1J1

121J1B

AH-800 AHRU NO. 1

E5E6

ARINC 429NO. 2 OUT

190J2A

C190J2A

G7G8L

ARINC 429 NO. 1 OUTH

K4K5

HL

ARINC 429NO. 1 IN

HL

ARINC 429NO 3. OUT

AZ-850 MADC NO. 19J1

6667

TO:STALL PROTECTION COMPUTER,GROUND PROXIMITYWARNING SYSTEM

AD-49869-R1@

IC-600 IAC NO. 12930

3334 L

LH

H

PRI AHRSARINC 429 IN

AP AHRSARINC 429 IN

IC-600 IAC NO. 2

8384

SEC AHRSARINC 429 INL

H

4950

HL

ARINC 429 AHRS NO. 4 IN

NZ-2000 FMS(OPTIONAL)

AB

FX-600 FLUX VALVENO. 1XY

CDEF

ZHC

STATOR

EXCITATION IN

STATOR COMMON

4J1

ORNREDBLKYELBRN

HL

ARINC 429IN NO 1

DA-800 DAU NO. 2137J1B

4243

F14F15L

ARINC 429 OUT NO. 4H

Figure 2-2-4. Pilot’s AHRS Interface



H3H4H7

H10

728

4443

K12K13

AIRCRAFTTESTCONNECTOR

28 V DC FROM DC BUS 2

DC GND28 V DC FROM BACKUP DC BUS2

HL

ARINC 429 AHRS NO. 5 IN

HL

ARINC 429NO. 3 OUT

SLVD

DG

AHRSTEST

H6

A10

CW

CCW

H1

H4

H2

K15

H3

COM012

WOW GROUND = ON GROUND

AB

FX-600 FLUX VALVENO. 2

H8H9H10H12H13H11

XY

CDEF

ZHC

H

L

F6F7F8F9

F11

INHIBITPWR

RETURNCLOCK

DATA

MEM ACCESS DATA TXMEM ACCESS DATA RX

MEM ACCESS COMMMEM ACCESS MODE SEL

28 V DC PRIMARY28 V DC AUXILARY

28 V DC RETURN

DG MODE SELECT

TEST COMMAND

HEADING SLEW

ABCDEF

FLUX VALVE INPUT

FLUX VALVEEXCITATION

FLUX VALVE COMMON

TOMEMORY MODULE

MEMORYMODULE

STATOR

EXCITATION IN

STATOR COMMON

A B CC1J1

121J1B

C4J1

AH-800 AHRU NO. 2

ARINC 429NO. 2 OUT

IC-600 IAC NO. 1

C190J2A

190J2A

L

ARINC 429NO. 1 OUT

H

SEC AHRSARINC 429 INL

H

K4K5

HL

ARINC 429NO. 1 IN

HL

ARINC 429NO 3. OUT

AZ-850 MADC NO. 2C9J1

6667

NZ-2000 FMS(OPTIONAL)

2223

HL

ARINC 429ATTITUDE IN

WU-660/880 RTA(WU-650/870 RTA)59J1

AD-49863@

IC-600 IAC NO. 2

PRI AHRSARINC 429 INL

HG7

G8

E5E6

29

30

83

84

ORNREDBLKYELBRN

(K)(L)

Figure 2-2-5. Copilot’s AHRS Interface



Table 2-2-4. AH-800 ARINC 429 Output Data

Parameter Label Units Data Range

Magnetic Heading 320 Degrees ±180 from North

Pitch Angle 324 Degrees ±180

Roll Angle 325 Degrees ±180

Pitch Rate 326 Deg/sec ±128

Roll Rate 327 Deg/sec ±128

Yaw Rate 330 Deg/sec ±128

Longitudinal Accel 331 G’s ±4

Lateral Accel 332 G’s ±4

Normal Accel 333 G’s ±4

Flux Valve Heading 334 Degrees ±180 from North

Raw Flux Valve Heading 100 Degrees ±180 from North



C. Modes of Operation

The AHRS has six operating modes:

• Initialization

• Full Performance

• DG

• Basic

• Test

• Maintenance.

(1) Initialization

Normal system initialization is performed on the ground. Upon application of power,the AH-800 performs self-test functions to determine the condition of its sensors,CPU, power supply, and I/O subsystems. During this time, the outputs are broughtinto alignment with the true local vertical and magnetic heading. This proceduretakes approximately 2 minutes on the ground and 15 seconds in the air. At thecompletion of the initialization sequence, all outputs are within their statedaccuracies and the system is in the full performance mode of operation, unless asystem input or flight crew command has placed the system in one of itsreversionary modes of operation.

The AH-800 will withstand power interruptions of 200 milliseconds or less, withoutthe loss of data. Prolonged power interruptions will require a restart during flight.Should this occur, all data will be flagged invalid until the system is reinitialized.This will take approximately 15 seconds to complete.

(2) Full Performance

The full performance mode is the systems normal operating configuration. While inthis mode, true airspeed is used in the computation of pitch and roll attitude toproduce a low gain, velocity damped erection loop and the flux valve is used as thelong-term heading reference. Refer to Table 2-2-5 for full performance accuracy.

(3) DG Mode

This mode allows the heading channel of the AHRU to operate as a free non-slavedgyro, which is not referenced to the flux valve. Entry into this mode is by pilotcommand (through the SLAVED/DG toggle switch) and is used when operating incharted areas of unreliable magnetic heading, or when a flux valve failure hasoccurred. The system will automatically synchronize to the flux valve heading, whenswitched from DG to slaved operation.

When in the DG operating mode, the DG may be slewed clockwise by moving theheading slew switch to the +360 position momentarily. Counterclockwise slewing isdone by moving the switch to the -360 position momentarily.



Table 2-2-5. AH-800 Full Performance Accuracy

Parameter Range Accuracy Resolution Free Drift Accuracy

Pitch AngularRate

0 to 70°/sec 0.05°/sec 0.05°/sec for ratesup to 25°/sec

0.5% for rates>25°/sec

<0.2%

Roll Angular Rate 0 to 70°/sec 0.05°/sec 0.05°/sec for ratesup to 25°/sec


<0.2%

Yaw Angular Rate 0 to 40°/sec 0.05°/sec 0.05°/sec for ratesup to 25°/sec


<0.2%

Pitch Attitude ±85° Static:<0.1°Dynamic:<0.5°

0.5° <1.5°

Roll Attitude ±360° Static:<0.1°Dynamic:<0.5°

0.5° <1.5°

Heading ±360° Static:<0.2°Dynamic:<1.0°

1.0° <5.0°/HR

Normal Accel ±5.0 g Static:<0.1 gDynamic:<0.1 g

±0.01 g <0.2%

Lateral Accel ±2.0 g Static:<0.1 gDynamic:<0.1 g

±0.01 g <0.2%

Long Accel ±2.0 g Static:<0.1 gDynamic:<0.1 g

±0.01 g <0.2%

(4) Basic Mode

This mode is entered automatically if the AHRU loses its True Airspeed (TAS) input.In this mode, the AHRU will not correct its pitch and roll computations foracceleration induced errors. This gives the AHRS the same accuracy as aconventional spinning mass type gyro.

(5) Test Mode

This mode can be activated by pilot command (through the AHRS TEST pushbutton)at any time while the aircraft is on the ground, with weight-on-wheels. During thetest mode, system outputs are driven to preset values to verify proper operation ofthe pitch, roll and heading channels, as well as interconnects and displays (refer toTable 2-2-6).



Table 2-2-6. AHRS ARINC 429 Output Test Mode Data

Label Signal Test Value

320 Mag Heading 030

324 Pitch Angle 15° up

325 Roll Angle 5° right

326 Pitch Rate 10°/sec

327 Roll Rate 10°/sec

330 Yaw Rate 10°/sec

331 Longitudinal Accel 0.02 g

332 Lateral Accel 0.1 g

333 Normal Accel 0.1 g

334 Flux Valve Heading 22.5°

(6) Maintenance Mode

This mode is designed for use in installing and maintaining the AHRS. The installerwould use this mode for the setting of discretes such as AHRU orientation andARINC 429 Hi/Low select, as well as determining flux valve compensationcoefficients. Honeywell Field Support uses this mode for special data retrieval onthe aircraft.

A special function available in the maintenance mode is the automatic compassswing. By using this feature, the time spent in initial installation, as well as byoperators performing routine compass system checks is reduced without requiringthe need for additional test equipment. Refer to SECTION 4 of this manual forAHRS compass swing information and procedures.

D. Reversionary Switching

Should an AHRS become invalid, the system can be re-configured to display the remainingvalid AHRS on both sides of the cockpit. Each crew member has a reversionary panelavailable. At power up, the on-side AHRS provides the display data for that side of thecockpit. Pressing the AHRS reversionary pushbutton causes the display electronics todisplay the cross-side AHRS data. When in this condition, the same AHRS is beingdisplayed on both sides. This condition is annunciated with an amber annunciation on bothPFDs.



4. Fault Monitoring

The AHRU contains monitors that ensure the accuracy and integrity of the system as follows:

• Ensure the accuracy of inertial calculations

• Ensure the availability of data on the digital buses

• Ensure accurate recording and retrieval of fault information

• Allow a system self-test to be initiated by operator command.

A. Fault Detection

The AHRS Built-In-Test (BITE) permits fault detection with a 90% minimum probability ofsuccess. It does this by performing power-on, continuous, pilot actuated, and groundmaintenance tests. The BITE is classified in three primary classes, according to thefollowing definition:

• Critical - Failures which would affect primary functions, causing erroneous outputs ofpitch and roll attitude, angular rates or accelerations.

• Non-critical - Failures not affecting primary functions, but can affect heading angle orheading angle rate.

• Maintenance - Unannunciated failures stored in BITE memory for maintenance useonly.



B. Power-On BITE

The power-on BITE tests are performed only in the power-on mode of operation. Power-ontests check the functions that cannot be tested during any of the other normal operationalmodes without interfering with normal operation. Power-on BITE will execute the following:

• Processor self-test

• Processor diagnostic test

• Processor program sumcheck test

• RAM addressing integrity test

• Dual port RAM addressing integrity test

• Discrete output wraparound test

• Accelerometer test

• Nonvolatile memory sumcheck test

• Memory module sumcheck test

• Auxiliary power and switchover test

• Calibration memory sumcheck test

• ARINC 429 transceiver VLSI wraparound test

• FOG EEPROM sumcheck test

• FOG electronics EEPROM sumcheck test.

NOTE: These tests are not performed during an in-air start.



C. Continuous BITE

The continuous tests are performed in all modes and do not interfere with normal operationof the AHRU. Continuous BITE will execute the following:

• CPU test

• Program EPROM sumcheck test

• RAM read after write test

• Power supply, reference, and excitation test

• Excessive power fluctuation test

• Stack overflow test

• Loop completion test

• Gyro hardover pulse count test

• IFOG optics test

• Accelerometer hardover reasonableness test

• Flux valve drive monitor test

• Flux valve open line monitor test

• Flux valve sine and cosine reasonableness test

• Sensor temperature test

• AHRU overtemperature test

• ARINC 429 transceiver VLSI wraparound test

• ARINC transmit test

• ARINC input test

• Controller transmit test

• Interrupt rate test

• ARINC bus recovery test

• Fault annunciator test.



D. Fault Reaction

Fault reaction for each class of BITE is described as follows.

(1) Critical

If a critical fault affecting the ARINC I/O buses has been detected, the followinghappens:

• All output data is flagged invalid

• A fault record is stored in nonvolatile memory.

(2) Non-Critical

A non-critical BITE fault is defined as a failure which does not affect the primaryfunctions (attitude, rate or acceleration), but can affect heading angle or headingangle rate. For these faults, the AHRS will revert to the non-slaved mode ofoperation and store the fault in NVM and invalidate the heading outputs on theARINC 429 data buses.

E. Maintenance

These faults do not immediately affect AHRS performance during this power cycle, but dorequire maintenance action. In this case, the AHRS will store the fault in NVM.

F. Fault Storage

The AHRS contains NVM for storage of BITE detected faults.

G. Flight Faults

The 200 most recent BITE faults will be stored in BITE memory in a zone designated forFlight Faults. The faults will be stored as records that contain the following data:

• Fault service status

• Miscellaneous status

• Fault type

• Fault event status

• Time since power-on

• Detailed fault code

• Snapshot parameters.



(1) Fault Service Status

This data is set by the maintenance operator to indicate that this stored fault hasbeen investigated and resolved. At the time of fault occurrence, the record is storedwith the status indicating Not Serviced.

(2) Miscellaneous Status

On-ground/In-air system status to be stored as part of the fault record. This data willalso contain an indication of power-up versus scheduled or continuous BITE and acounter for repeat occurrences of this fault during this power cycle.

(3) Fault Type

This identifies the fault at the system sub-component level from among the following:

• AHRU fault

• Flux valve interface fault

• Memory module interface fault

• Controller interface fault

• No ARINC 429 data

• AHRU overtemperature fault.

H. Ground Faults

If the AHRS determines that it is being operated on the ground and that the aircraftengines are inactive, then BITE detected faults will be stored in an area of BITE memoryseparate from flight fault storage. The ground fault storage can hold the last 24 detectedBITE failures stored in a form similar to those of the flight fault variety previouslydiscussed.



I. Fault Indications

Fault indications are presented on the PFD, MFD, and EICAS display tubes. See Figures2-2-6 through 2-2-8.

(1) PFD

See Figure 2-2-6.

AHRS faults that are displayed on the PFD are as follows.

(a) Loss of valid pitch or roll information:

• Removal of the pitch tape

• Removal of the roll pointer

• Removal of the flight director bars

• Entire attitude sphere is cyan

• Red ATT FAIL annunciation in the top half of the attitude sphere

• Inhibit of the attitude miscompare annunciation.

(b) Loss of valid heading information:

• Red HDG FAIL annunciation displayed in the top of compass arc

• Removal of heading bug and digital readout

• Removal of the course/desired track pointer

• Amber dash of the course/desired track digital readout

• Removal of the drift bug

• Removal of the lateral deviation pointer

• Removal of the TO/FROM display

• Inhibit of the heading miscompare annunciation

• Removal of the absolute bearing pointers

• Amber dash of the current heading digital readout in arc mode.



AD-51250@

N

S

33

15

3012

WE

246

21

3

- - - CRS

HDG- - -

0

1

2

3

3

1

2

ATTFAIL

ADC2SG2

ATT2

3000

LNAV VS VNAVAP YD ASEL

14500

29.92 IN

0020

80

14500

143

14000

9

1

260

280

220

200

240

260

.410 M

2

R1

HDGFAIL

1000


Figure 2-2-6. PFD AHRS Failure Indications



(2) MFD

See Figures 2-2-7 and 2-2-8.

AHRS faults that are displayed on the MFD with loss of valid heading information areas follows.

(a) MAP MODE

• Removal of the tick mark labels

• Amber dash of the heading digital readout

• Red HDG FAIL annunciation displayed in the top of the compass arc

• Removal of the heading bug and digital readout

• Removal of the drift bug

• Removal of all waypoint symbols

• Removal of all NAVAID symbols

• Removal of all airport symbols

• Removal of holding pattern racetrack symbol

• Removal of the lateral deviation display

• Removal of the designator symbol, bearing/distance readout, and LAT/LONreadout.

(b) PLAN MODE

• Removal of the aircraft symbol.





INVALID CURRENTHEADING

(3 AMBER DASHES)

HDGFAIL

HEADING FAILANNUNCIATOR

(RED)

N

S

33

15

3012

WE

246

21

3

+15

300

SATTATTAS

FMSKDVT12.512

NMMIN

50 50


AD-51251@

STABTGTWX-16

+25

Figure 2-2-7. MFD MAP MODE AHRS Failure Indications




25

+15+25300

SATTATTAS

FMSLLO112.512

NMMIN

TGTTX


1200 1200N

AD-51252@


HDGFAIL

HEADING FAILANNUNCIATOR

(RED)

Figure 2-2-8. MFD PLAN MODE AHRS Failure Indications



(3) EICAS

AHRS conditions that generate messages on the CAS display and their meaningsare as follows.

(a) AHRS 1-2 OVERHEAT

The AHRS fan located on the mounting tray has failed, or the fan filter is dirtyand needs to be cleaned.

(b) AHRS 1-2 BASIC

The AHRU operating mode has changed to BASIC due to a loss of trueairspeed which is required for acceleration induced error compensation.



(Blank Page)





2.3 ADZ-850 MICRO AIR DATA SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-1


A. AZ-850 Micro Air Data Computer (MADC) . . . . . . . . . . . . . . . . . . . . 2-3-2

B. BL-870 PFD Bezel Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-4

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-5

A. Pilot’s Air Data System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-5

B. Copilot’s Air Data System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-7

C. Modes of Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-9

D. MADC Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-9

E. Static Source Error Correction (SSEC) . . . . . . . . . . . . . . . . . . . . . . 2-3-9

F. Operational Range . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-10

G. Overspeed Warning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-10


A. Primary Flight Display (PFD) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-13

B. Multifunction Display (MFD) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-15

C. EICAS Display (EICAS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-16




FIGURE/TITLE PAGE


Figure 2-3-1. AZ-850 Micro Air Data Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-2

Figure 2-3-2. BL-870 PFD Bezel Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-4

Figure 2-3-3. Pilot’s MADC Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-6

Figure 2-3-4. Copilot’s MADC Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-8

Figure 2-3-5. Vmo/Mmo Curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-10

Figure 2-3-6. PFD MADC Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-14

Figure 2-3-7. MFD MAP MODE MADC Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . 2-3-15

List of Tables

TABLE/TITLE PAGE

Table 2-3-1. AZ-850 Micro Air Data Computer Leading Particulars . . . . . . . . . . . . . . . . . . 2-3-3

Table 2-3-2. BL-870 PFD Bezel Controller Leading Particulars . . . . . . . . . . . . . . . . . . . . . 2-3-4

Table 2-3-3. AZ-850 MADC Performance Accuracy . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-11

Table 2-3-4. MADC Functional Test Outputs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3-12



SECTION 2.3ADZ-850 MICRO AIR DATA SYSTEM

1. General

The PRIMUS® 1000 Integrated Avionics System installed in the Embraer 145 aircraft includestwo AZ-850 Micro Air Data Computers (MADC). Each MADC contains the sensing units andassociated electronics in a small and efficient package. The MADC takes inputs of static airpressure and pitot pressure, total air temperature and baro set information, performs thenecessary computations, and transmits air data information via ARINC 429.

The MADC also includes Static Source Error Correction (SSEC), and outputs various discretesfor airspeed, Mach, vertical speed, overspeed warning, etc.

Barometric correction is input to the MADC directly from a rotary set knob located on the on-side PFD bezel. Air data target values are displayed as digital quantities and are shown asmoving bugs on the PFD air data displays.

The system is comprised of the following LRUs:

• AZ-850 Micro Air Data Computer (MADC)

• BL-870 PFD Bezel Controller.




A. AZ-850 Micro Air Data Computer (MADC)

Figure 2-3-1 shows a graphical view of the AZ-850 Micro Air Data Computer (MADC). TheMADCs are located in the avionics nose bay of the aircraft. Table 2-3-1 provides itemsand specifications that are particular to the computer.

AD-22798-R1@

Figure 2-3-1. AZ-850 Micro Air Data Computer



Table 2-3-1. AZ-850 Micro Air Data Computer Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.04 in. (102.62 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.04 in. (153.42 mm)

• Length (from rear connector) . . . . . . . . . . . . . . . . 5.71 in. (145.03 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.5 lb. (2.05 kg)

Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . 28 V dc, 16 Watts (maximum)

User Replaceable Parts . . . . . . . . . . . . . . . . . . . . . . . None

Mating Connectors:

• Static, straight . . . . . . . . . . . . . . . . . . . . . . . . . . MS24393-6 and Nut MS24400-6

• Static, elbow . . . . . . . . . . . . . . . . . . . . . . . . . . . MS24394-6 and Nut MS24400-6

• Pitot, straight . . . . . . . . . . . . . . . . . . . . . . . . . . . MS24393-4 and Nut MS24400-4

• Pitot, elbow . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS24394-4 and Nut MS24400-4

• Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS27473E20B-35S

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MT-840 Tray, HPN 7014702-901

Pitot pressure and static pressure are accepted by the MADC through MS-type threadedfittings on the front panel. All electrical input/output is through a multi-pin connector on thefront panel. The MADC is mounted in a tray, as shown in Figure 2-?, and the tray has noelectrical or pneumatic connectors.

The following air data values are output by the AZ-850 MADC:


• Pressure altitude

• Indicated airspeed (IAS)

• Mach number

• Vertical speed (VS)

• Maximum operating speed (Vmo)

• Static and total air temperature (SAT and TAT)

• True airspeed (TAS).



B. BL-870 PFD Bezel Controller

Figure 2-3-2 shows a graphical view of the BL-870 PFD Bezel Controller. They are locatedas part of the PFDs on the instrument panel. Table 2-3-2 provides items and specificationsthat are particular to the controller.

IN/HPA STD

BARO

AD-50627@

Figure 2-3-2. BL-870 PFD Bezel Controller

Table 2-3-2. BL-870 PFD Bezel Controller Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.33 in. (3.37 cm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.71 in. (17.05 cm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.73 in. (4.33 cm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.3 lb (0.135 kg)


• Knob, BARO . . . . . . . . . . . . . . . . . . . . . . . . . . . HPN 7000895-3

• Setscrew . . . . . . . . . . . . . . . . . . . . . . . . . . . . .(Hex Socket, 6-32 x 3/16", Cup point)

HPN 0455-224

As part of the air data system, the BL-870 PFD Bezel Controller contains the BARO setknob, STD pushbutton, and IN/HPA pushbutton. The BARO knob and STD pushbuttoninterface directly with the on-side MADC. The IN/HPA pushbutton interfaces directly withboth DC-550 Display Controllers.



3. Operation

A. Pilot’s Air Data System

Figure 2-3-3 shows the Pilot’s MADC Interface. The inputs to the MADC consist of staticand pitot pressure, total air temperature, baro set information, aircraft configuration data,weight-on-wheels discrete, and a functional test discrete. The pressure inputs are bymeans of MS-type threaded fittings. The total air temperature input is a 500 ohmARINC 575 total air temperature probe.

Baro correction is controlled by the BARO knob mounted on the BL-870 PFD bezelcontroller. The Baro Set range is from 16.00 to 32.00 inHg, or from 542 to 1084 hPa.Selection of inHg or hPa is done through the IN/HPA PFD bezel pushbutton. Rotation ofthe BARO set knob sends grey code pulses directly to the on-side MADC, which providethe MADC with barometric correction information, so as to output altitude above Mean SeaLevel (MSL). Clockwise rotation increases inHg in 0.01 increments, or hPa in 1.0increments. Counterclockwise rotation of the knob decreases the selected setting by a likeamount.

The STD (standard) pushbutton adjacent to the BARO set knob commands the MADC toset the barometric correction to 29.921 inHg or 1013.24 hPa.

Aircraft configuration data is supplied by seven discrete inputs. Each input is defined asopen or grounded for the type of aircraft. This ensures that the proper MADC is installed,and sets the aircraft dependent parameters of Vmo and SSEC.

Side select data is supplied by three discrete inputs. Each input is defined as open orgrounded for the side of the aircraft in which the MADC is installed. This ensures that theMADC receives and transmits the appropriate side air data.

The ADC1 or ADC2 test switch located on the maintenance panel behind the pilot’s seat isused for functional test.

The WOW input is used to inhibit test during flight.

The ARINC 429 No. 1 output sends air data information to both IAC No. 1 and IAC No. 2.

The ARINC 429 No. 2 output sends air data information to channel A of FADEC No. 1 andNo. 2 and to the horizontal stabilizer control unit.

The ARINC 429 No. 3 output sends airspeed information to the AHRU No. 1 and altitudeinformation to COM Unit No. 1 and No. 2.

The ARINC 429 No. 4 output sends airspeed and altitude to the GPWS, stall protectioncomputer, pressure digital controller and optional FMS and GNSSU.



12

2627

HL

28 V DC INPUT PWR

WOW

DC GROUND

28 V DC FROMAM ESS DC BUS 1

55TO MAINTENANCE TESTCONNECTOR 56

57

RS-232 DATA RCVRRS-232 COMMONDATA XMTR

70

ADC 1 TEST SWITCH

HL

6869

ARINC 429NO. 4 OUT

7574

71

HL

BAROCORRECTION

BARO STD SYNC

3231 H

L

STATIC PRESSURE INPUT

PITOT PRESSURE INPUT

500 OHMTEMPERATURE

PROBE

AZ-850 MADC NO. 1 9J1

9J1

46

2125

130J1

H

L

BAROSET

STD BARO

DU-870 NO. 1 (PFD)

121J1BNZ-2000 FMS

6768

ARINC 429 MADC NO. 12 IN

HL

TEST

FUNCTIONALTEST

15

TEMPPROBE

PROBE HEATSIGNAL

UNIT ID 3UNIT ID 5SDI NO. 1SIGNAL GROUND

358

13

SIGNALGROUND

HL

6364

FADEC 1A, FADEC 2A, ANDHORIZONTAL STABILIZERCONTROL UNIT

ARINC 429NO. 2 OUT

21 AIRSPEED SWITCH137 KNOTS

TO RUDDERTRIM SYSTEM

28OVERSPEEDWARNING

TO CREW ALERTINGSYSTEM

(OPTIONAL)

TO:

STALL PROTECTION COMPUTER,PRESSURE DIGITAL CONTROLLER,GROUND PROXIMITY WARNINGSYSTEM

190J2BIC-600 IAC NO. 1

2324

PRI MADCARINC 429 IN

HL

C190J2AIC-600 IAC NO. 2

4344

SEC MADCARINC 429 IN

HL

143J1RCZ-851 COM UNIT NO. 1

9986


HL

C143J1RCZ-851 COM UNIT NO. 2

10069


HL

6061

ARINC 429NO. 1 OUT

ARINC 429NO. 3 OUT

6667

1J1BAH-800 AHRU NO. 1

K4K5

MADCARINC 429 IN

HL

HL

TO:

HL

16BARO SEL(In/HPa)

GNSSU

67

ARINC 429MADC NO.1 IN

HL

(OPTIONAL)

82 BARO SEL OUT

149J1

GROUND = AIRCRAFT ON GROUND

ALTITUDE SWITCH24,500 FEET

22

23

AIRSPEED SWITCH205 KNOTSTO SPOILER

COMMANDSYSTEM

AD-49862-R1@

9J1

Figure 2-3-3. Pilot’s MADC Interface



B. Copilot’s Air Data System

Figure 2-3-4 shows the Copilot’s MADC Interface. The copilot’s system is identical to thepilot’s system with the following exceptions:

• The BARO knob and the STD pushbutton are both located on the copilot’s BL-870 PFDBezel Controller

• The ARINC 429 No. 1 output connects to IAC No. 1 secondary MADC input and to IACNo. 2 primary MADC input

• The ARINC 429 No. 2 output connects to FADEC 1B and 2B input and the HorizontalStabilizer Control Unit

• The ARINC 429 No. 3 output connects to the No. 2 AHRU and RCZ-851 COM Unit No.2 primary input and RCZ-851 COM Unit No. 1 secondary input. It is also connected tothe optional PRIMUS® 660/870/880 Weather Radar Air Data input.

• The ARINC 429 No. 4 output connects to the optional FMS MADC No. 3 input, theoptional GNSSU MADC No. 2 input and to the Stall Protection Computer, PressureDigital Controller, and the Windshear Computer.



12

26

27

HL 28 V DC INPUT PWR

WOW

DC GROUND

28 V DC FROMAM DC BUS 2A

55TO MAINTENANCE TESTCONNECTOR 56

57

RS-232 DATA RCVRRS-232 COMMONDATA XMTR

70

ADC 2 TEST SWITCH

7574

71

HLBARO CORRECTION

BARO STD SYNC

3231 H

L

STATIC PRESSURE INPUT

PITOT PRESSURE INPUT

500 OHMTEMPERATURE

PROBE

AZ-850 MADC NO. 2 C9J1

C9J1

46

2125

HL

BAROSET

STD BARO

DU-870 NO. 5 (PFD)

59J1 WU-660/880 RTA(WU-870 RTA)

4344

ARINC 429AIR DATA IN

HL

TEST

FUNCTIONALTEST

15

TEMPPROBE

PROBE HEAT

UNIT ID 3UNIT ID 5SDI NO. 2SIGNAL GROUND

35913

SIGNALGROUND H

L6364

FADEC 1B, FADEC 2B, ANDHORIZONTAL STABILIZERCONTROL UNIT

ARINC 429NO. 2 OUT

21 AIRSPEED SWITCH137 KNOTS

TO RUDDERTRIM SYSTEM

28OVERSPEEDWARNING

TO CREW ALERTINGSYSTEM

TO:STALL PROTECTION COMPUTERPRESSURE DIGITAL CONTROLLER,WINDSHEAR COMPUTER

C190J2BIC-600 IAC NO. 2

2324


HL

190J2AIC-600 IAC NO. 1

4344


HL

C143J1RCZ-851 COM UNIT NO. 2

9986


HL

143J1RCZ-851 COM-UNIT NO. 1

10069


HL

6061

ARINC 429NO. 1 OUT

ARINC 429NO. 3 OUT

6667

C1J1BAH-800 AHRU NO. 2

K4K5

MADCARINC 429 IN

HL

HL

TO:

HL

16BARO SEL(In/HPa)

GNSSU10

11

ARINC 429MADC NO.2 IN

HL

(OPTIONAL)

82 BARO SEL OUT

149J1


AIRSPEED SWITCH205 KNOTS

22TO SPOILER

COMMANDSYSTEM ALTITUDE SWITCH

24,5OO FEET22

HL

6869

ARINC 429NO. 4 OUT

NZ-2000 FMS

9076

ARINC 429 MADC NO. 13 IN

HL

(OPTIONAL)121J1A

AD-49861-R1@

C9J1

C130J1

(t)(r)

Figure 2-3-4. Copilot’s MADC Interface



C. Modes of Operation

The AZ-850 MADC has three operating modes: normal, initiated test mode andmaintenance test mode. The normal mode is for normal aircraft operation.

The initiated test mode is activated by setting the applicable ADC test switch located on themaintenance panel to test. This test is interlocked with the weight-on-wheels switch and isinhibited when airspeed is greater than 50 knots. In the initiated test mode, the MADCoutputs are driven to preset values to check the operation of the MADC, interconnects anddisplays.

The Maintenance Test Mode provides the capability to display maintenance pages on thePFD while on the ground (WOW).

See Tables 2-3-3 and 2-3-4 for performance accuracy and MADC self test output data.

D. MADC Monitoring

A nonvolatile memory is provided for the on the ground analysis of any in flight monitortrips. This memory is accessed through the aircraft test connector.

Built in monitoring routines include tests to ensure that:

• All program memory is addressable and readable

• Pressure sensor outputs are in the correct range

• The aircraft electrical keying is correct

• Power supply outputs are of the correct values

• The inputs to the MADC are reasonable

• The central processing unit is functioning properly.

E. Static Source Error Correction (SSEC)

SSEC refers to a correction to account for errors which are long term, measurable andrepeatable. Typical static sensing systems are built as flush openings in the side of theaircraft, or as a protruding probe. The airflow past the static port will cause the pressure inthe static system to be different from the undisturbed air.

In general terms, a pressure error is caused by anything which causes a variation in theairflow as it passes the static port. Examples are:

• Airflow around the curved fuselage

• Landing gear extended

• Position of flaps

• Aircraft yawing motion

• Angle of attack changes.



F. Operational Range

The AZ-850 MADC is capable of providing airspeed and altitude data over the rangesspecified below:

• Altitude: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -2,000 to 60,000 feet

• Calibrated Airspeed: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30 to 500 knots

• True Airspeed: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50 to 599 knots.

G. Overspeed Warning

During the normal mode of operation, overspeed warning is a signal to the pilot when theCAS output has exceeded the current value of Vmo. Overspeed warning function in thefunctional test mode is to illustrate that the warning signal operates correctly.

The AZ-850 MADC computes the overspeed warning using calibrated airspeed, Mach,Vmo, Mmo, and pressure altitude. Overspeed warning is computed by comparing thepressure altitude with the break point in the Vmo/Mmo curve. See Figure 2-3-5. The breakpoint is the point at which Vmo is constant or linear below and Mmo is constant or linearabove. If pressure altitude is at or below the break point in the curve, then the CAS will becompared to Vmo to switch overspeed warning on and off. If pressure altitude is above thebreak point in the curve, then Mach will be compared to Mmo to switch overspeed warningon and off.

SS

EC

CO

RR

EC

TE

D P

RE

SS

UR

E A

LT

ITU

DE

(F

EE

T)

CALIBRATED AIRSPEED (KNOTS)

37,000

26,268

10,000

8,000

250 320

.780 M

AD-51311-R1@

Figure 2-3-5. Vmo/Mmo Curve



Table 2-3-3. AZ-850 MADC Performance Accuracy

Parameter Range Accuracy SSec

Barometric Altitude -2,000 to 60,000 ft From -2,000 ft to sea level: . . ±15 ftFrom sea level to 20,000 ft: . . ±20 ftFrom 20,000 ft to 30,000 ft: . . ±40 ftFrom 30,000 ft to 50,000 ft: . . ±80 ftFrom 50,000 ft to 60,000 ft: . ±150 ft

Yes

Altitude Rate ±20,000 fpm ±30 fpm or ±5% (largest) Yes

Mach 0 to 1.0 Variable from ±0.003 to ±0.050 Mdepending on speed and altitude

Yes

Computed Airspeed 30 to 500 kts From 30 to 60 kts: . . . . . . . . ±5 ktsAt 80 kts: . . . . . . . . . . . . . . ±3 ktsFrom 100 to 200 kts: . . . . . . ±2 ktsAt 500 kts: . . . . . . . . . . . . . ±5 kts

Yes

True Airspeed 50 to 599 kts At 50 kts . . . . . . . . . . . . . . ±12 ktsAt 70 kts . . . . . . . . . . . . . . . ±8 ktsFrom 150 to 599 kts . . . . . . ±4 kts

Yes

Static Air Temperature -99 to + 60 °C ±1 °C No

Total Air Temperature -78 to +99 °C ±1 °C No

Baro Corrected Altitude -2,000 to 60,000 ft ±5 ft or 0.5% (largest) No

Baro Correction 542 to 1084 hPa16.0 to 320 inHg

± 1.0 hPa± 0.01 inHg

No



Table 2-3-4. MADC Functional Test Outputs

PARAMETER ARINC 429 VALUE

Pressure Altitude 4,000 ft

Baro Corrected Altitude 1,000 ft

Altitude Rate 5,000 ft/min

Calibrated Airspeed 325 kts

True Airspeed 325 kts

Mach 0.77

Static Air Temperature -45 °C

Total Air Temperature -16 °C

Baro Correction inHg 29.921 inHg

Baro Correction mB 1013.0 mB

Static Pressure 29.92 inHg

Total Pressure 1013.2 mB

Impact Pressure 181.8 mB

Vmo 320 kts

Vmo Warning Active for 2 seconds

MADC Valid Inactive

Output Discretes Active



4. Fault Monitoring

Fault indications are presented on the primary flight, multifunction, and engine instrumentdisplays.

A. Primary Flight Display (PFD)

Figure 2-3-6 shows the fault indications as presented on the PFD. MADC faults displayedon the PFD are as follows.

(1) If barometric altitude is invalid:

• The digits, ticks, and chevrons on the altitude tape are removed and a red X isplaced over the tape

• The selected altitude digits are replaced with five amber dashes

• The selected altitude bug is removed.

(2) If indicated airspeed is invalid:

• The digits and ticks on the airspeed tape are removed and a red X is placed overthe tape

• The airspeed trend vector is removed

• The Vspeed bugs, if selected, are removed

• The selected airspeed bug and digital readout is removed.

(3) If vertical speed (altitude rate) is invalid:

• The vertical speed pointer and digital readout are removed

• The vertical speed target bug and digital display are removed

• A boxed VS is displayed vertically in the center of the vertical speed arc.



AD-51253@

- - - - - - - -

.- - - M

CRS

NM

HDG

0

1

2

3

3

1

2

- -.- - IN

- - - -

IAS

ALT

VS

10 10

10 10

20 20

2250

N

S

33

15

3012

WE

246

21

3

TGT

200 RA

LOC HDG IAS GSAP YD

GS

360

13.1

360

VOR1

ADF2


Figure 2-3-6. PFD MADC Failure Indications



B. Multifunction Display (MFD)

Figure 2-3-7 shows the fault indications as presented on the MFD. MADC faults displayedon the MFD are the same for the MAP and PLAN formats.

(1) If TAS is invalid, the digits are replaced with three amber dashes.

(2) If temperature is invalid, the SAT and TAT digits are replaced with three amberdashes.



INVALID TAT(3 AMBER DASHES)

INVALID TAS(3 AMBER DASHES)

N

S33

15

30

12

W

E

24

6

21

3

25

SATTATTAS

FMSKDVT12.512

NMMIN

360

50 50

31547.0


AD-51254@

STABTGTWX-16

INVALID SAT(3 AMBER DASHES)

Figure 2-3-7. MFD MAP MODE MADC Failure Indications



C. EICAS Display (EICAS)

MADC fault indications are displayed on the EICAS in the form of an E1-2 ADC DATA FAILadvisory message. The FAIL message is enabled whenever the temperature probe fails orthere is a loss of data from the MADC.





2.4 AA-300 RADIO ALTIMETER SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-1


3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-4

A. Auxiliary Radio Altitude Output . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-4

B. Primary Radio Altitude Output . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-4

C. RA Minimum Annunciation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-4

D. Low Altitude Awareness . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-5

E. Radio Altitude Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-5

4. Fault Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-8




FIGURE/TITLE PAGE


Figure 2-4-1. RT-300 Radio Altimeter Receiver Transmitter . . . . . . . . . . . . . . . . . . . . . . . 2-4-2

Figure 2-4-2. Single AA-300 Radio Altimeter Interface Diagram . . . . . . . . . . . . . . . . . . . . 2-4-6

Figure 2-4-3. Optional Dual AA-300 Radio Altimeter Interface Diagram . . . . . . . . . . . . . . 2-4-7

Figure 2-4-4. PFD Radio Altitude Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4-8

List of Tables

TABLE/TITLE PAGE

Table 2-4-1. RT-300 Radio Altimeter Receiver Transmitter Leading Particulars . . . . . . . . 2-4-3



SECTION 2.4AA-300 RADIO ALTIMETER SYSTEM

1. General

The PRIMUS® 1000 Integrated Avionics System installed in the Embraer 145 aircraft includes asingle AA-300 Radio Altimeter System as standard. A second system is available as a customerpurchased option. The AA-300 Radio Altimeter System is a dual antenna, short pulse systemdesigned to provide the flight crew with absolute altitude above the terrain information. Therange of the AA-300 is from 0 to 2500 feet. The system is designed for automatic continuousoperation over wide variations of terrain, weather, target reflectivity and aircraft attitude.

Radio altitude information is displayed on the lower portion of the ADI sphere on the pilot’s andcopilot’s PFDs. If the system becomes invalid, an amber RA is annunciated instead of the digitson the PFDs to alert the pilots.

The system also generates preset altitude trip outputs for other aircraft systems. These outputssupply a ground potential at the preset altitudes of 50, 200, 1200, and 1500 feet. The AA-300Radio Altimeter System is also the trigger to display the decision height (MIN) annunciation onthe PFDs and is also used to gain program the localizer and glideslope signals in the flightdirector.

The system is comprised of the following LRUs:

• A single RT-300 Radio Altimeter Receiver Transmitter (RT)

• Two Radio Altimeter Antennas (non-Honeywell).




Figure 2-4-1 shows a graphical view of the RT-300 Radio Altimeter Receiver Transmitter. It islocated in the avionics nose bay of the aircraft. Table 2-4-1 provides items and specificationsthat are particular to the unit.

AD-2092@

ZEROALTITUDEADJUSTMENT

Figure 2-4-1. RT-300 Radio Altimeter Receiver Transmitter

The RT contains the necessary power supplies, radio frequency transmitting and receivingcircuitry and timing circuits with which to determine the aircraft absolute altitude. The systemtransmits radio frequency pulses, measures the time until the reflection is received anddetermines the aircraft absolute altitude above terrain.

It generates a DC output voltage which is proportional to the absolute altitude. It can alsogenerate preset altitude trip outputs for other aircraft systems. These outputs supply a groundpotential at or below the preset altitudes of 50, 200, 1200, and 1500 feet.



Table 2-4-1. RT-300 Radio Altimeter Receiver Transmitter Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.09 in. (104.0 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.56 in. (115.8 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.07 in. (281.2 mm)

Weight (maximum) . . . . . . . . . . . . . . . . . . . . . . . . . 4.6 lb. (2.05kg)

Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . 28V dc, 17 watts

Transmitter Characteristics (nominal):

• Type . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Short-pulse modulation

• Pulse Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100 nanoseconds

• PRF . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Jittered between 12.5 and 50 kHz

• Radio Frequency (RF) . . . . . . . . . . . . . . . . . . . . . 4.3 GHz

• Peak Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5 W

Receiver Characteristics (nominal):

• Type . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Superheterodyne

• Intermediate Frequency (IF) . . . . . . . . . . . . . . . . . 60 MHz

Operational Altitude . . . . . . . . . . . . . . . . . . . . . . . . . 0 - 2,500 ft

Data Outputs/Accuracy:

• Precision Output . . . . . . . . . . . . . . . . . . . . . . . . . DC analog voltage (0 - 2,500 ft)

Gradient: -4.0 mV dc/ft0 alt = 0 volt

Accuracy: 0 - 100 ft, ±3 ft100 - 500 ft, ±3%500 - 2500 ft, ±4%

• Auxiliary Output . . . . . . . . . . . . . . . . . . . . . . . . . . DC analog voltage (0 - 2500 ft)

Gradient: Per ARINC 552• For Alt < 480 ft, Output is:

0.02 (Alt) + 0.4• For Alt > 480 ft, Output is:

[10 + 10 ln] [(Alt+20)/500]

Accuracy: 0 - 100 ft, ±4%100 - 500 ft, ±4%500 - 2,500 ft, ±5%



Table 2-4-1. RT-300 Radio Altimeter Receiver Transmitter Leading Particulars

Item Specification

User Replaceable Parts . . . . . . . . . . . . . . . . . . . . . . None

Mating Connectors:

• J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS3126F16-26S

• J2 - TRANSMIT . . . . . . . . . . . . . . . . . . . . . . . . . . TNC - Male(Straight) GRFF 4007-0002

• J3 - RECEIVE . . . . . . . . . . . . . . . . . . . . . . . . . . . TNC - Male(Straight) GRFF 4007-0002

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hard Mount

3. Operation

See Figures 2-4-2 and 2-4-3 for AA-300 Radio Altimeter System interface information.

The RT-300 Receiver Transmitter outputs an analog voltage that represents aircraft absolutealtitude above the terrain. This primary output signal is used for the following PFD displays:

• Radio altitude

• DH annunciation

• Radio altitude low altitude awareness.

The primary analog voltage is also used in the flight director for gain programming of thelocalizer and glideslope signals.

A. Auxiliary Radio Altitude Output

The system generates an ARINC 552 DC analog output that is sent to the TCAS ComputerUnit for use in controlling TCAS operational constraints between 1 and 1,450 feet AGL.

B. Primary Radio Altitude Output

The system generates a -4.0 mV dc/foot output voltage which is proportional to theaircraft’s absolute altitude. This output is sent to both Integrated Avionics Computers(IACs) where it is used by the Electronic Display System (EDS) and other subsystems. Ina dual AA-300 Radio Altimeter System, RT No. 1 sends its outputs to IAC No. 1 and RTNo. 2 sends its outputs to IAC No. 2.

C. RA Minimum Annunciation

The pilot can set a radio altitude minimum value based on RA through the use of the RAknob on the DC-550 Display Controller. When the aircraft is at or below the RA minimumset value, the annunciation MIN will appear in the lower right side of the ADI sphere. TheMIN annunciation is amber inside a white window.



D. Low Altitude Awareness

Radio altitude low altitude awareness is provided for display on the barometric altitudetape. The presentation of low altitude awareness is used as an indication of the ground,with respect to the current barometric altitude.

When radio altitude is less than 550 feet, the lower portion of the altitude tape will start tochange color to brown. A horizontal line will be drawn across the altitude tape at thetransition. The shaded brown portion of the tape will rise in a linear fashion for radioaltitudes between 0 and 550 feet. At 0 radio altitude, the entire lower portion of the altitudetape will be brown.

Any of the conditions that follow will cause the low altitude awareness raster and horizontalline to be removed:

• Radio Altitude indicates invalid

• Barometric Altitude indicates invalid.

E. Radio Altitude Test

The radio altimeter test is functional when the TST button is activated on the DC-550Display Controller. When the test is active, the following is seen on the PFD:

• Radio altitude digital readout of 100 feet

• Low altitude awareness is displayed on the barometric altitude tape.

The radio altimeter can be tested on the ground or in the air. The test is inhibited if theflight director is in the glideslope capture mode or windshear mode.



RT-300 RADIOALTIMETER R/T NO.1

POWER GND

TRIP COMMON

TEST SWITCH

PRIMARY OUTPUT(H)

OUTPUT TEST

OUTPUT COMMON(C)

20J1

P

T

X

N

E

RA VALID IN

RAD ALTINPUT

+

-

28 V DC INPUT PWR28 V DC FROMAM DC BUS 1A

c

b

IC-600 IAC NO.1190J2A

IC-600 IAC NO.2C190J2A

RAD ALTINPUT

+

-

RT-910 TCASCOMPUTER UNIT

RAD ALT NO.1VALID IN

RAD ALT NO.1INPUT

+-

193RMP

W

RA VALID OUT Y

AUX OUTPUT

TRIP NO. 2 (1500 FT)

TRIP NO. 4 (200 FT)

V

L

STALL PROTECTIONCOMPUTER

AURAL WARNINGSYSTEM

49

5150

49

5150

2J

2H

RAD ALT NO.2 VALID IN

RAD ALT NO.2INPUT

+-

193RBP

2K

3A3B

3C

AD-49875@

97

RA VALID IN

RA TEST OUT

97 RA TEST OUT

Figure 2-4-2. Single AA-300 Radio Altimeter Interface Diagram



RT-300 RADIO ALTIMETER R/T NO.1

POWER GND

TRIP COMMON

TEST SWITCH

PRIMARY OUTPUT(H)

OUTPUT TEST

RA VALID OUT

OUTPUT COMMON(C)

20J1

P

X

W

N

E

Y

RA VALID IN

RAD ALTINPUT

+

-

28 V DC INPUT PWR28 V DC FROMAM DC BUS 1Ac

b

IC-600 IAC NO.1

97 RA TEST OUT

RT-300 RADIOALTIMETER R/T NO.2

POWER GND

TRIP COMMON

TEST SWITCH

PRIMARY OUTPUT(H)

OUTPUT TEST

RA VALID OUT

OUTPUT COMMON(C)

C20J1

P

T

W

N

E

Y

28 V DC PWR c

bIC-600 IAC NO.2

RA TEST OUT

RA VALID IN

97

C190J2A

49

51 RAD ALTINPUT

+

-

RT-910 TCASCOMPUTER UNIT (OPTIONAL)


RAD ALT NO.1INPUT

+

-

193RMP

2H

2J

2K


RAD ALT NO.2INPUT

+

-

193RBP

3A

3B

3C

T

X

AUX OUTPUT (H)

V

L STALL PROTECTIONCOMPUTER

AURAL WARNINGSYSTEM

190J2A

49

51

TRIP NO. 2 (1500 FT)

TRIP NO. 4 (200 FT)

AUX OUTPUT (H)

V

L STALL PROTECTIONCOMPUTER

AURAL WARNINGSYSTEMTRIP NO. 2 (1500 FT)

TRIP NO. 4 (200 FT)

28 V DC FROMAM DC BUS 2A

AD-49874@

50

50

Figure 2-4-3. Optional Dual AA-300 Radio Altimeter Interface Diagram



4. Fault Monitoring

Figure 2-4-4 shows fault indications. With the loss of valid radio altitude information, thefollowing is indicated on the PFD:

• The radio altitude digital readout is removed

• A red RA is displayed in the attitude sphere

• The decision height indications are inhibited

• The low altitude awareness raster is removed.

1400

29.92 IN

CRS

ADF2

HDG

AD-51255-R1@

10 10

10 10

2020

360

330

33

LOC GS

AP

0

1

2

3

3

1

2

500

-RA-

RADIOALTITUDEINVALIDINDICATION


.410 M

ILS1

13.1 NM

200 RA

1500

500

Figure 2-4-4. PFD Radio Altitude Failure Indications





2.5 WEATHER RADAR SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-1


A. Weather Radar Receiver Transmitter Antenna (RTA) . . . . . . . . . . . . 2-5-2

B. Weather Radar Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-6

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-9

A. Target Alert (TGT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-9

B. Rain Echo Attenuation Compensation Technique (RCT) Mode . . . . 2-5-15

C. Turbulence (TRB) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-15

D. Test (TST) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-15

E. Weather (WX) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-15

F. Ground Map (GMAP) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-15

G. Standby (SBY) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-16


A. PFD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-17

B. MFD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-18

C. PRIMUS® 650/870 Weather Radar Test Mode . . . . . . . . . . . . . . . . . 2-5-19

D. PRIMUS® 660/880 Weather Radar Test Mode . . . . . . . . . . . . . . . . . 2-5-21

(1) On-Ground TEST Display (with TEXT FAULTS Enabled) . . . 2-5-21

(2) In Flight TEST Display (with TEXT FAULTS Enabled) . . . . . 2-5-22

(3) Fault Monitors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-22




FIGURE/TITLE PAGE


Figure 2-5-1. Typical Weather Radar Receiver Transmitter Antenna . . . . . . . . . . . . . . . . 2-5-2

Figure 2-5-2. Maximum Permissible Exposure Level Boundary . . . . . . . . . . . . . . . . . . . . 2-5-5

Figure 2-5-3. Optional Weather Radar Controllers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-6



Figure 2-5-6. PFD Weather Radar Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-17

Figure 2-5-7. MFD Weather Radar Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-18

Figure 2-5-8. PRIMUS® 650/870 MFD Weather Radar Test Mode Indications . . . . . . . . . 2-5-19

Figure 2-5-9. PRIMUS® 660/880 MFD Weather Radar Test Mode Indications . . . . . . . . . 2-5-23

List of Tables

TABLE/TITLE PAGE

Table 2-5-1. Weather Radar Receiver Transmitter Antenna Leading Particulars . . . . . . . . 2-5-3

Table 2-5-2. Weather Radar Controller Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . 2-5-7

Table 2-5-3. WC-XXX Control Functions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-7

Table 2-5-4. Target Alert Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-5-9





SECTION 2.5WEATHER RADAR SYSTEM

1. General

The PRIMUS® 1000 Integrated Avionics System includes the PRIMUS® 650, 660, 870 or 880Weather Radar system, which consists of one WU-XXX Receiver Transmitter Antenna (RTA),one WC-XXX Weather Radar Controller, and parts of the EDS. Each system is an X-Banddigital radar designed for weather location and analysis, and for ground mapping.

The radar system detects precipitation in storms along the flight path of the aircraft and givesthe flight crew a visual indication, in color, of storm intensity and turbulence. In the weatherdetection mode, target returns are displayed at one of five video levels (0, 1, 2, 3, or 4), with 0represented by a black screen because of weak or no returns, and levels 1, 2, 3, and 4represented by green, yellow, red, and magenta respectively, to show progressively strongerreturns. Areas of potentially hazardous turbulence are shown in gray white. In ground−mappingmode, video levels of increasing reflectivity are displayed as black, cyan, yellow, and magenta.

The radar information may be displayed on the Multifunction Displays (MFDs) and on thePrimary Flight Displays (PFDs). The radar range, radar operating mode, and antenna tiltfunctions are all controlled by the WC-XXX Weather Radar Controller.




A. Weather Radar Receiver Transmitter Antenna (RTA)

Figure 2-5-1 shows a graphical view of a typical Weather Radar Receiver TransmitterAntenna (RTA). It is located in the nose of the aircraft. Table 2-5-1 provides items andspecifications that are particular to the unit.

AD-14251, SH1@

Figure 2-5-1. Typical Weather Radar Receiver Transmitter Antenna



Table 2-5-1. Weather Radar Receiver Transmitter Antenna Leading Particulars

Item Specification


• Base Diameter . . . . . . . . . . . . . . . . . . . . . . . 10.04 in. (255.0 mm)

• Height (Antenna flat) . . . . . . . . . . . . . . . . . . 10.06 in. (225.5 mm)

• Height (Antenna full arc . . . . . . . . . . . . . . . . 16.04 in. (407.4 mm)

Weight (maximum):

• WU-650 . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.7 lb (6.22 kg)

• WU-870 . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16.0 lb (7.26 kg)

• WU-660/880 . . . . . . . . . . . . . . . . . . . . . . . . 15.7 lb (7.12 kg)

Primary Power . . . . . . . . . . . . . . . . . . . . . . . . . . +22 to +32 V dc, 110 W (max)

Antenna:

• Size . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12-inch flat plate radiator

• Stabilization . . . . . . . . . . . . . . . . . . . . . . . . . Line-of-sight, ±30 degrees

• Tilt . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±15 degrees

• Scan (Full) . . . . . . . . . . . . . . . . . . . . . . . . . 120 degrees (±60 degrees)

• Scan (Sector) . . . . . . . . . . . . . . . . . . . . . . . 60 degrees (±30 degrees)

Transmitter (WU-650/870):

• Frequency . . . . . . . . . . . . . . . . . . . . . . . . . . 9345 ± 25 MHz

• Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.3 kW (nom), magnetron

• Pulse Width . . . . . . . . . . . . . . . . . . . . . . . . . 1.2, 1.5, 2.4, 4.8, 9, 18, and 27 µSecdetermined by selected range and mode

• PRF . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 120, 240, 360, and 480 Hz, determined byselected range in WX 480/1260 in Turb, and840 in GCR

Transmitter (WU-660/880):

• Frequency . . . . . . . . . . . . . . . . . . . . . . . . . . 9375 ± 25 MHz

• Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 kW (nom), magnetron

• Pulse Width . . . . . . . . . . . . . . . . . . . . . . . . . 1.0 and 2.0 µSec (nom), determined byselected range and mode

• PRF . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 120, 240, and 420 Hz, determined byselected range and mode



Table 2-5-1. Weather Radar Receiver Transmitter Antenna Leading Particulars

Item Specification

Receiver (WU-650/870):

• Noise Figure . . . . . . . . . . . . . . . . . . . . . . . . 8.5 dB, typical

• Intermediate Frequency (IF) . . . . . . . . . . . . . 35 MHz

• IF Bandwidth . . . . . . . . . . . . . . . . . . . . . . . . 0.8 MHz (nom)

• Video Bandwidth . . . . . . . . . . . . . . . . . . . . . Commensurate with selected pulse width

• STC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Present in all modes

• MDS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -112.4 dBm (nom) on 300 NM range

Receiver (WU-660/880):

• Intermediate Frequency (IF) . . . . . . . . . . . . . 60 MHz, 1st conversion;10.7 MHz, 2nd conversion

• IF Bandwidth . . . . . . . . . . . . . . . . . . . . . . . . 0.725 MHz (nom)

• Video Bandwidth . . . . . . . . . . . . . . . . . . . . . Commensurate with selected pulse width

• STC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Present in all modes

• MDS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -115 dBm (nom) on 300 NM range

Displayed Ranges:

• WX/MAP . . . . . . . . . . . . . . . . . . . . . . . . . . . 5, 10, 25, 50, 100, 200, and 300 NM fullscale

• Flight Plan . . . . . . . . . . . . . . . . . . . . . . . . . . 5, 10, 25, 100, 200, 300, 500, and 1000 NMfull scale

Mating Connector:

• WU-650 . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS3126F20-41S, HPN 4000809-606

• WU-870 . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS3126F22-55S, HPN 4000809-626

• WU-660/880 . . . . . . . . . . . . . . . . . . . . . . . . Glenair PN DD104F1000, HPN 7517883-3



The WU-XXX RTA is an integrated unit which incorporates transmitter, receiver, andantenna into a single unit. The RTA uses an 12−inch flat plate radiator. The remainder ofthe circuitry is contained in the electronics package which forms the RTA base. The RTAtransmits and receives X−band radio frequency energy for the purposes of weatherdetection and ground mapping. The transmitted signals are sent directly to the antennafrom the transmitter circuitry which is mounted on the rear of the antenna. Echo signalsreceived by the antenna are applied directly to the receiver. The receiver and processingsystem processes these signals by encoding them into one of four levels depending ontheir intensity, scan converts them and outputs the scan converted data to the DU-870Display Units No. 1, 2, 4, and 5.

Heating and radiation effects of weather radar can be hazardous to life. Personnel shouldremain at a distance greater than "R" (see Figure 2-5-2) from the radiating antenna inorder to be outside of the envelope in which radiation exposure levels equal or exceed10 mW/cm2, the limit recommended in FAA Advisory Circular AC No. 20−68B, August 8,1980, Subject: "Recommended Radiation Safety Precautions for Ground Operation ofAirborne Weather Radar." The radius R, to the Maximum Permissible Exposure Level(MPEL) boundary is calculated for the radar system of the basis of radiator diameter, ratedpeak−power output, and duty cycle. The greater of the distances calculated for either thefar−field or near−field is based on the recommendations outlined in AC No. 20−68B.

IEEE Standard for Safety Level with Respect to Human Exposure to Radio FrequencyElectromagnetic Fields 3 kHz to 300 GHz (IEEE C95.1-1991), recommends an exposurelevel of no more than 6 mW/cm2.

Honeywell, Inc. recommends that operators follow a 5 mW/cm2 standard. Figure 2-5-2shows MPEL for 10 mW/cm2 and 5mW/cm2 exposure levels.

RADARSYSTEM

PRIMUS 650/870

â

PRIMUS 660/880

â 12 IN30.5 CM

16.5 FT5.0M

25.7FT7.8 M

AIRCRAFTNOSE ANDRADOME

AIRCRAFT LUBBER LINE

RADIUSR

MPELBOUNDARY

270°

DIAMETEROF FLAT-

PLATERADIATOR

5 mW/CM ² RADIUS ROF MPEL

BOUNDARY

10 mW/CM ² RADIUS ROF MPEL

BOUNDARY

12 IN45.7 CM

6 FT1.84 M

9 FT2.75 M

AD-51256@

Figure 2-5-2. Maximum Permissible Exposure Level Boundary



B. Weather Radar Controller

Figure 2-5-3 shows a graphical view of various Weather Radar Controllers. It is located onthe pedestal in the cockpit of the aircraft. Table 2-5-2 provides items and specificationsthat are particular to the unit.

SECTTGTSTABRCT

GAIN RADAR SLV TILT

+

-MIN MAX

PULLVAR

OFF

SBYWX GMAP

FP

TST

OFF

SECTTGTGCRTRB

GAIN RADAR SLV TILT

+

-MIN MAX

PULLVAR

OFF

SBYWX RCT

GMAP

FP

TST

PULLAUTO

SECTTGTSTABTRB

GAIN RADAR SLV TILT

+

-MIN MAX

PULLVAR

OFF

SBYWX RCT

GMAP

FP

TST

PULLACT

OFF

WC-650/660 CONTROLLERUSED ON PRIMUS 650/660 RADAR INSTALLATIONSR

WC-870 CONTROLLERUSED ON PRIMUS 870 RADAR INSTALLATIONSR

WC-880 CONTROLLERUSED ON PRIMUS 880 RADAR INSTALLATIONSR AD-51249@

Figure 2-5-3. Optional Weather Radar Controllers



Table 2-5-2. Weather Radar Controller Leading Particulars

Item Specification


• Length (from rear of bezel) . . . . . . . . . . . . . . . . 7.0 in. (177.8 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.75 in. (146.1 mm)

• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.87 in. (47.5 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.0 lb (0.86 kg)

Power Requirements:

• Primary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +22 to +32 V dc, 8.5 watts, maximum

• Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5 V ac/dc at 1.0 A, nominal

User Replaceable Parts . . . . . . . . . . . . . . . . . . . . . . None

Knobs:

• GAIN and TILT Knob . . . . . . . . . . . . . . . . . . . . HPN 7011875-901

• RADAR Mode Knob . . . . . . . . . . . . . . . . . . . . . HPN 7011875-902

Setscrews (Multi-Spline, 2-56 by 1/8 inch, Cup Point) . HPN 2500148-64

Mating Connector . . . . . . . . . . . . . . . . . . . . . . . . . . MS27473F14-18S

NOTE: The mating connector uses strain relief MS27506-B14-2.

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Unit Dzus Fasteners

The WC-XXX Weather Radar Controller sets the modes, and antenna tilt used to displayradar information on the PFDs and MFD. The function of each switch and control is listedin Table 2-5-3.

Table 2-5-3. WC-XXX Control Functions

Item Functional Description

TILT Single-turn rotary control that varies antenna tilt between 15 degrees upand 15 degrees down. The range between +5 degrees and -5 degrees isexpanded for ease of setability.• AUTO TILT (PULL) - Places elevation control under Auto Tilt which

adjusts antenna tilt in relation to altitude and selected range. Tiltknob can be used for fixed offset corrections of up to ± 2.0 degrees.Available on WU-870 RTA only.

• ACT TILT (PULL) - Places elevation control under AltitudeCompensated Tilt (ACT) which adjusts antenna tilt in relation toaltitude and selected range. Tilt knob can be used for fixed offsetcorrections of up to ± 2.0 degrees. Available on WU-880 RTA only.



Table 2-5-3. WC-XXX Control Functions

Item Functional Description

RADARMODE

A six-position (WC-650/660) or seven-position (WC-870/880) rotary switchthat selects the following primary radar modes:• OFF - Removes power from system.• SBY - Standby. Places system in non-transmitting mode.• WX - Selects the system in the weather (WX) mode.• RCT - (WU-650/660 only) Enables the cyan (sky blue) REACT (RCT)

field to indicate ranges at which the receiver calibration has beenexceeded in the WX mode. RCT is selected in the TST mode onalternate sweeps automatically. RCT compensation is active in allmodes except GMAP.

• GMAP - Places system in ground map (GMAP) mode. RCTcompensation is inactive.

• FP - Selects the system flight plan (FP) (navigation) display mode.• TST - Selects the system self-test mode.

GAIN Single-turn rotary control that varies the RTA receiver gain. Control isactive when pulled. When pushed, receiver gain is preset and calibrated.Selection of RCT or TGT ALERT overrides the variable gain settingcausing receiver gain to be fixed and calibrated. When active (pulled),full ccw rotation provides straight gain increase over the preset value.

TRB Momentary button that selects the turbulence (TRB) mode. In this mode,areas of potentially hazardous turbulence are displayed in gray-whitecolor, in addition to normal reflectivity data.TRB may only be engaged in the WX mode and in selected ranges of 50NM or less. This function is not available in the WU-650/660 RTA.

STAB Momentary button that deselects the stabilization function. Also, used toinvoke the stabilization trim mode. This function is not available in theWU-870 RTA. On WC-660/880 controllers, the STAB button is used tooverride Forced Standby by pushing it four times within 3 seconds.

TGT Momentary alternate-action button that selects the target alert (TGT)function (OFF RANGE TGT ALERT).

SECT Momentary alternate-action button that selects either full (120 degrees) orreduced (60 degrees) scan sector (SCT).

RCTWC-650/660only

Momentary button enables the REACT (RCT). RCT is always selected inthe TEST mode. RCT compensation is available in all modes exceptMAP.

Range A two-button range selection function that permits range selection from 5to 300 NM full scale in WX, RCT, or GMAP mode or 5 to 1000 NM fullscale in the Flight Plan mode. The up arrow button selects increasingrange while the down arrow button selects decreasing ranges.On WC-650/870 controller, these buttons are also used to exit ForcedStandby by pressing both simultaneously.



3. Operation

See Figures 2-5-4 and 2-5-5 for Weather Radar System interface information.

The RTA accepts mode, tilt, etc., commands from the Weather Radar Controller on a serialcontrol bus. The RTA outputs mode, range, tilt, etc., commands to the IACs on two EDS controlbuses, and outputs scan converted data to the DUs on two EDS picture buses. Antennastabilization data is input from the No. 2 AHRU, in ARINC 429 format. True airspeed for theturbulence processors is input from the No. 2 MADC, in ARINC 429 format. REACTcompensation override is grounded so that selection of RCT mode on the Weather RadarController overrides the GAIN control setting and forces preset gain.

The following paragraphs describe each of the modes and features controlled by the WC-XXXradar controller.

A. Target Alert (TGT)

The TGT button allows the pilot to select or deselect the target alert mode of the radarsystem. Target alert is selectable in any WX range except 300 NM. The target alert circuitmonitors for red level or greater targets within ±7.5 degrees of dead ahead. Also the targetmust have the following depth and range characteristics as shown in Table 2-5-4.

Table 2-5-4. Target Alert Characteristics

SelectedRange (NM)

TargetDepth (NM)

TargetRange (NM)

5 5 5 - 55

10 5 10 - 60

25 5 25 - 75

50 5 50 - 100

100 5 100 - 150

200 5 200 - 250

300 Disabled

FP (Flight Plan) 5 5 - 55

It should be noted that while target alert is functional at the above ranges, it is improbablethat a realistic target would be strong enough to be detected if its range exceeds five timesthe displayed range. Also, note that the target alert is inactive within the displayed range.Selecting target alert prevents variable gain from being selected.



(Blank Page)



IC-600 IAC NO. 1

WU-650/870 RECEIVER/TRANSMITTER/ANTENNA

K

L

59J159J1

LEFT EFISCONTROL BUS

H

L

17

18

LEFT EFISCONTROLBUS

H

L

m

H

L

61J1

C190J2A

AH-800 AHRU NO. 2

H

L

K12

K13

190J2A

C1J1B

C130J1

LEFTPICTURE BUS

DU-870 NO. 1 (PFD)

H

L

77

78

LEFT EFISPICTUREBUS

H

L

130J1

LEFTPICTURE BUS

DU-870 NO. 2 (MFD)

H

L

77

78

131J1

ARINC 429NO. 3 OUT

T GND ENABLES 429STABILIZATION (WU-650 ONLY)

ARINC 429ATTITUDE INPUT

C131J1

RIGHT EFISCONTROLBUS

H

L

e

c

REMOTE ON

NO. 1CONTROL BUS

H

L

RIGHT EFISPICTURE

BUS

H

L

k

s

H

L

28V DC INPUT PWR

AB

XW

R REACT COMPENSATIONOVERIDE (GND/OPEN)

AD-49859@

64TERMINATION

28 V DC FROMAM DC BUS 1B


AIRCRAFTLIGHTING BUS

RIGHT EFISCONTROL BUS

IC-600 IAC NO. 2

H

L

17

18

WC-650/870RADAR CONTROLLER

PRIMARYCONTROL BUS

H

L

REMOTE ON

A

B

R

RIGHTPICTURE BUS

DU-870 NO. 5 (PFD)

H

L

77

64

RIGHTPICTURE BUS

DU-870 NO. 4 (MFD)

H

L

77

78

TERMINATION

78

D 28V DC INPUT POWER

P WOW INPUT

H

E

F CHASSIS GROUND

5V LIGHTING INPUT

LIGHTING GROUND

POWER GROUND

ANNUN DIMMING

J

K

d

U

f

FROM PEDESTALOFF/BRT CONTROL

n

g

h

H ARINC 429 ADC (NO)(WU-870 ONLY)

T WOW INPUT(WU-870 ONLY)

WOWSW

IN AIR

ON GROUND

H

L

AZ-850 MADC NO. 2

H

L

66

67

C9J1

ARINC 429NO. 3 OUT

(WU-870 INSTALLATIONS ONLY)

ARINC 429AIR DATA IN(WU-870 ONLY)

t

r

Figure 2-5-4. PRIMUS® 650/870 Weather Radar System Interface

Page 2-5-11/1222-05-14 Nov 1/1996Use or disclosure of information on this page is subject to the restrictions on the title page of this document.


IC-600 IAC NO. 1

WU-660/880 RECEIVER/TRANSMITTER/ ANTENNA

22

23

59J159J1

LEFT EFISCONTROL BUS

H

L

17

18

LEFT EFISCONTROLBUS

H

L

30

H

L

61J1

C190J2A

AH-800 AHRU NO. 2

H

L

K12

K13

190J2A

C1J1B

C130J1

LEFTPICTURE BUS

DU-870 NO. 1 (PFD)

H

L

77

78

LEFT EFISPICTUREBUS

H

L

1

130J1

LEFTPICTURE BUS

DU-870 NO. 2 (MFD)

H

L

77

78

131J1

ARINC 429NO. 3 OUT

ARINC 429ATTITUDE IN

C131J1

RIGHT EFISCONTROLBUS

H

L

32

71

REMOTE ON

NO. 1CONTROL BUS

H

L

RIGHT EFISPICTURE

BUS

H

L

5

6

H

L

28V DC INPUT PWR

1314

1516

AD-50546@

64TERMINATION

28 V DC FROMAM DC BUS 1B


AIRCRAFTLIGHTING BUS

RIGHT EFISCONTROL BUS

IC-600 IAC NO. 2

H

L

17

18

WC-660/880RADAR CONTROLLER

PRIMARYCONTROL BUS

H

L

REMOTE ON

A

B

R

RIGHTPICTURE BUS

DU-870 NO. 5 (PFD)

H

L

77

64

RIGHTPICTURE BUS

DU-870 NO. 4 (MFD)

H

L

77

78

TERMINATION

78

D 28V DC INPUT POWER

P WOW INPUT

H

E

F CHASSIS GROUND

5V LIGHTING INPUT

LIGHTING GROUND

POWER GROUND

ANNUN DIMMING

J

K

31

2

43

44

AZ-850 MADC NO. 2

H

L

66

67

C9J1

ARINC 429NO. 3 OUT

H

L

ARINC 429AIR DATA IN

72

11

33

FROM PEDESTALOFF/BRT CONTROL

85 ALT/AIR/STAB CONFIG(GND/OPEN)

95 TEXT FAULTS DISCRETE(GND/OPEN)

90 WOW INPUT(WU-880 ONLY)

AIRCRAFTWOW

SWITCH

IN AIR

ON GROUND

Figure 2-5-5. PRIMUS® 660/880 Weather Radar System Interface



B. Rain Echo Attenuation Compensation Technique (RCT) Mode

The rain echo attenuation compensation technique (REACT) function permits the radarreceiver to adjust its own sensitivity automatically to compensate for attenuation losses asthe radar pulse passes through weather targets on its way to illuminate other targets. Thisis done by measuring the intensity of signals, and deducing from them the density, andtherefore, the attenuation of the target, and then using this information to adjust thesensitivity. This is done continuously on each radar azimuth radial. There is a maximumvalue to which sensitivity may be set due to the receiver generating noise, and would fillthe display with noise if it were too high. When this maximum value is reached, a bluecolor is displayed for the remainder of the displayed range. This gives the pilot anunmistakable warning that attenuation is hiding possible severe weather areas that cannotbe accurately detected. REACT is always selected in TEST mode. REACT is available inall modes except GMAP.

C. Turbulence (TRB)

When the turbulence submode is selected (WU-870/880 RTAs only), the radar processesreturn signals in order to determine if a turbulence signature is present. Areas ofpotentially hazardous turbulence are displayed as gray white. The high power of thePRIMUS® 870/880 permits detection of hazardous turbulence in areas of otherwise weaklyreflective rainfall. Any areas shown as turbulence should be avoided. TRB may only beengaged in the WX mode and in selected ranges of 50 NM or less.

D. Test (TST)

Used to select a special test pattern to allow verification of system operation; 100-milerange is automatically selected; TEST is displayed in mode field. Transmitter output poweris radiated in TEST mode. Any faults present will be displayed when selecting TESTmode. See Table 2-5-5.

E. Weather (WX)

Used to select weather detection operation. If selected prior to end of the warmup period,WAIT will be displayed until the transmitter warms up (approximately 50 to 90 seconds).WX is displayed in mode field. Transmitter output power is radiated in the WX mode. Inthe WX mode, four precipitation levels are displayed as green, yellow, red, and magenta.

F. Ground Map (GMAP)

The ground-mapping operation is selected by setting the mode control to GMAP. The TILTcontrol is turned down until the desired amount of terrain is displayed. The degree ofdown-tilt will depend upon the aircraft altitude and the selected range. The receiver STCcharacteristics are altered to provide equalization of ground-target reflection versus range.As a result, the selection of preset GAIN will generally provide the desired mappingdisplay. However, the pilot may desire to decrease the gain manually by selecting manualgain and rotating the GAIN control. Transmitter output power is radiated in GMAP mode.



G. Standby (SBY)

Standby is useful for keeping radar in ready state while taxiing, loading, etc. In standby,the antenna does not scan, the transmitter is disabled, the display memory is erased, andthe antenna is stowed in a tilt-up position.

Standby should be selected anytime it is desired to keep power on the system withouttransmitting.



4. Fault Monitoring

Fault indications are presented on the PFD and MFD displays. See Figures 2-5-6 thru 2-?.

A. PFD

If a weather radar failure occurs, regardless of the display format (FULL/ARC/WX), anamber WX will be annunciated on the lower left side of the HSI. See Figure 2-5-6.

10 10

10 10

20 20

9

1

260

280

220

200

240

260 14500

.410M

360 CRS

VOR1

ADF2

HDG001

0

1

2

3

3

1

2

1000

29.92 IN

LNAV VSAP YD

3000

0020

80

14500

143

14000

2

R1

N33 3

WX 50

360MAG2

AD-51257@NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS.

2o

FAILED WX(AMBER)

20 20

Figure 2-5-6. PFD Weather Radar Failure Indications



B. MFD

If a weather radar failure occurs, an amber WX is displayed in the WX mode annunciationdisplay field. See Figure 2-5-7.

N

S

33

15

30

12

W

E

24

6

21

3

25

+15+25300

SATTATTAS

FMSKDVT12.512

NMMIN

360

50 50

PLAB1

PLAB2

*PBD01LL01

KDVT

2

MAPPLANWXTCASCKLSTMFDSYS

TCAS TESTABVFL

MAG2

RA NO BRGTA NO BRG

-05

-10

12

00

31546.0

MENU

6

AD-51258@NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS.

WX

FAILED WX(AMBER)

Figure 2-5-7. MFD Weather Radar Failure Indications



C. PRIMUS® 650/870 Weather Radar Test Mode

After an amber WX is annunciated, the fault codes can be read by entering into the TESTmode. Figure 2-5-8 illustrates the PRIMUS® 650/870 MFD Weather Radar Test Mode, andTable 2-5-5 lists the PRIMUS® 650/870 Fault Codes and Fault Descriptions.



360

N

3

6

3330

5050

FAIL25

TEST MODEFAIL

ANNUNCIATOR

FAULT CODE

TEST PATTERN

25

+15

300

SATTATTAS

FMSKDVT12.512

NMMIN

31547.0


AD-51259-R1@

+25

Figure 2-5-8. PRIMUS® 650/870 MFD Weather Radar Test Mode Indications



Table 2-5-5. PRIMUS® 650/870 Fault Codes

Fault Code Fault Description

03 Analog to Digital Converter Failure, > 1 minute

04 DADC Low Speed 429 Failure for > 4 minutes

06 Airspeed (DADC or Analog) > 700 kts for > 4 minutes

07 Pulse Pair Processor Failure for > 1 minute

13 +15 Volts Failure (> ± 12.5 V or > 17.5 V for > 1 minute)

14 Parallel Altitude change > 1000 feet

16 Magnetron Voltage Failure (< 1500 Volts, > 2700 Volts, > 1 minute)

21 Azimuth scanning incorrectly (> 2.5 degrees for > 1 minute)

22 STAB reference (< 1/2 A/D scale for > 4 minutes)

23 Automatic Gain Control Failure (< -1.1 V or > 10.5 V for > 1 minute)

24 Mixer Current Failure for > 1 minute

25 AFC Lock Failure for > 1 minute or 10 unlocks < 1 minute

26 Fan Voltage Abnormal for > 4 minutes

27 VLSI Failure-Loss of VALID READY Interrupt

30 Non-Volatile Memory Failure

31 Antenna Elevation error (> 2 degrees for > 1 minute)

32 NAV Computer High Speed ARINC 429 Failure for > 4 minutes

33 -15 Volts Failure (> - 18 V or > -12.5 V for > 1 minute)

35 AFC Sweep Failure ( <1.2 V or > 12.7 V for > 1 minute)

36 Altitude Failure - If input is > 60,000 feet for > 4 minutes

37 RAM Test Failure



D. PRIMUS® 660/880 Weather Radar Test Mode

The PRIMUS® 660/880 Weather Radar Systems can provide fault information on one oftwo formats:

• Fault codes

• Text codes.

The selection is made during initial system installation. All Embraer 145 aircraft arestrapped to display text faults.

When a fault occurs (either in-flight or on-ground), an amber WX overrides the modeannunciation on the PFD and/or MFD displays. The fault annunciation remains until thefault condition clears.

If fault annunciation occurs, the flight crew should select the test mode and note thedisplayed text fault and take appropriate action

Figure 2-5-9 illustrates the PRIMUS® 660/880 MFD Weather Radar Test mode and Table2-5-6 lists the PRIMUS® 660/880 Fault Codes and Fault Descriptions.

(1) On-Ground TEST Display (with TEXT FAULTS Enabled)

When TEST is initiated on the ground (Weight-On-Wheels asserted), six fields aredisplayed as shown in Figure 2-5-9. The six fields are as follows:

• Pilot Message Field (e.g. "STAB UNCAL")

• Line Maintenance Message (e.g. "CHK ATT SRC")

• Fault Code/Power-On Count (e.g. CODE:27 POC:0)

• Fault Name (e.g. "NO STAB SRC")

• Xmit On/Off (e.g. "XMIT ON!")

• Strap Code (e.g. "1F1BB:STRAPS").

Faults (up to 32) from the last 10 power-on cycles are cycled every two antennasweeps (approximately 8 seconds). That is, a fault will be displayed if and only if itoccurred within the last 10 power-on cycles and it is among the 32 most recent faultsto have occurred.

POC=0 is the current power-on cycle, POC=1 is the last power-on cycle, -2 is 2power-on cycles ago, etc.

Upon entering TEST mode, if there are no currently active faults, a "RADAR OK"message will be displayed for one sweep. After that the most recent fault isdisplayed, cycling to the oldest fault in the eligible list of faults. Upon reaching thelast fault, an "END OF LIST" message will be displayed. To recycle through the listagain, exit and re-enter the TEST mode.

Input-type faults (NO STAB SRC, NO AIRSPEED, NO ALTITUDE, etc.) will bedisplayed, but not logged, on-ground.



(2) In Flight TEST Display (with TEXT FAULTS Enabled)

NOTE: The radar will be transmitting when TEST is initiated while in theair.

If the Weight-On-Wheels input is not asserted, only the fields that follow aredisplayed:

• Pilot Message Field (e.g. "RADAR UNCAL")

• Line Maintenance Message (e.g. "PULL RTA")

• Fault Code.

Only currently active (not cleared) faults are displayed in-flight.

(3) Fault Monitors

Critical functions in the RTA are continuously monitored. Each fault condition has acorresponding 2-digit Fault Code (FC). Additionally, a FAULT NAME, a PILOTMESSAGE, and a LINE MAINTENANCE MESSAGE are associated with each faultcondition. These are given in Table 2-5-6.

The FAULT NAME describes what fault has been detected.

The PILOT MESSAGE advises the flight crew how to respond to a fault when itoccurs in-flight. This may include checking other systems, or to use caution wheninterpreting certain data displayed, and/or to advise that a minor function such asACT is unavailable.

The LINE MAINTENANCE MESSAGE advises the ground crew on a suggestedaction to take, or which LRU(s) to suspect.

The XREF code is a four-bit central aircraft information maintenance system (CAIMS)fault code. This code is not visible to the flight crew. It is used only on aircraft witha CAIMS installed.





360

N

3

6

33

30

5050

TEST-10

TEST MODEANNUNCIATOR

TEST PATTERNWITH TEXT

FAULT SHOWN XM

IT O

N! 1F1BB:STRAPSCO

DE: 27 POC 0 NO STAB SR

CSTA

B UNCAL CHK ATT SR

C

25

+15

300

SATTATTAS

FMSKDVT12.512

NMMIN

31547.0


AD-51619@

+25

Figure 2-5-9. PRIMUS® 660/880 MFD Weather Radar Test Mode Indications




FC XREF FAULTDESCRIPTION

FAULT NAME PILOTMSG

LINEMAINT

FAULT TYPE

4808 STARTUP CODECRC

4809 IOP CODE CRC

1 4810 DSP CODE CRC FLASH CRC RADARFAIL

PULL RTA POWER ON

4904 CONFIG TABLECRC

4905 FPGA FIRMWARECRC

4846 2V ADCREFERENCE

CONTINUOUS

4903 IOP READY

2 4908 INT ARINC 429LOOPBACK

IOP RADARFAIL

PULL RTA POWER ON

4913 ARINC 429 INTCOUPLING

4806 EEPROM TIMERCRC

FLASH CRC RADARFAIL

PULL RTA

3 4842 STAB TRIM CRC EEPROM REDOSTABTRIM

REDOSTABTRIM

POWER ON

4912 CALIBRATIONCRC

IOP RADARFAIL

PULL RTA

4812 IOP MAILBOX

4 4818 DSP MAILBOX MAILBOXRAM

RADARFAIL

PULL RTA POWER ON

4813 TIMING FPGARAM

4814 TIMING FPGAREG

5 4815 IO FPGA RAM FPGA RADARFAIL

PULL RTA POWER ON

4828 FPGADOWNLOAD

4906 IO FPGA REG





FAULT NAME PILOTMSG

LINEMAINT

FAULT TYPE

6 4847 STC MONITOR STC DAC RADARFAIL

PULL RTA POWER ON

7 4830 HVPS MONITOR HVPS MON RADARFAIL

PULL RTA CONTINUOUS

4816 DSP RAM

4817 DSP VIDEO RAM POWER ON

4855 DSP WATCHDOG CONTINUOUS

10 4900 MAILBOXMISCOMPARE

DSP RADARFAIL

PULL RTA

4901 DSP HOLDAASSERTED

POWER ON

4902 DSP HOLDA NOTASSERTED

4825 FILAMENTMONITOR

11 4827 SEVEREMAGNETRON

MAGNETRON RADARFAIL

PULL RTA LATCHED

4829 PFN TRIMMONITOR

HVPS MON CONTINUOUS

13 4832 ELEVATIONERROR

EL POSITION TILTUNCAL

CHKRADOME/

RTA

CONTINUOUS

14 4833 AZIMUTH ERROR AZ POSITION TILTUNCAL

CHKRADOME/

RTA

CONTINUOUS

15 4836 OVER TEMP OVER TEMP RADARCAUTION

PULL RTA CONTINUOUS

16 4837 XMITTER POWER XMTRPOWER

RADARUNCAL

PULL RTA CONTINUOUS

4839 NO SCI CONTROL

20 4911 NO ARINC 429CONTROL

NO CNTL IN CHKCNTLSRC

CHKCNTLSRC

PROBE





FAULT NAME PILOTMSG

LINEMAINT

FAULT TYPE

4840 AGC LIMITING PICTUREUNCAL

CONTINUOUS

21 4927 AGC RX DACMONITOR

AGC PULL RTA

4928 AGC TX DACMONITOR

RADARFAIL

POWER ON

22 4841 SELFTEST OSCFAILURE

RCVR SELFTEST

PICTUREUNCAL

PULL RTA CONTINUOUS

4843 MULTIPLE AFCUNLOCKS

4845 AFC SWEEPING SPOKINGLIKELY

CONTINUOUS

24 4929 AFC RX DACMONITOR

AFC PULL RTA

4930 AFC TRIM DACMONITOR

RADARFAIL

POWER ON

4848 AHRS/IRSSOURCE

27 4852 ANALOG STABREF

NO STABSRC

STABUNCAL

CHK ATTSRC

INSTALLATION

4849 DADC AIRSPEED CHK ADC

30 4932 AHRS/IRSGROUNDSPEED

NOAIRSPEED

TURBUNCAL

CHKSPEED

SRC

INSTALLATION

4931 DADC ALTITUDE CHK ADC

33 4933 AHRS/IRSINERTIALALTITUDE

NO ALTITUDE NO ACT CHK ALTSRC

INSTALLATION

34 4853 SCAN SWITCHOFF

SCANSWITCH

SCANSWITCH

CHKSWITCH

INSTALLATION

35 4854 XMIT SWITCHOFF

XMIT SWITCH XMITSWITCH

CHKSWITCH

INSTALLATION





FAULT NAME PILOTMSG

LINEMAINT

FAULT TYPE

4914 INVALIDALTITUDE/

AIRSPEED/STABSTRAPPING

36 4915 INVALIDCONTROLLER

SOURCESTRAPPING

INVALIDSTRAPS

RADARUNCAL

CHKSTRAPS

POWER ON

4916 CONFIG1DATABASE

VERSION/SIZEMISMATCH

IOP RADARFAIL

PULL RTA



(Blank Page)





2.6 SRZ-850 INTEGRATED RADIO SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-1


A. RM-855 Radio Management Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-2

(1) RMU Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-6

(2) Backup Navigation Display . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-8

(3) Backup Engine Display . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-20


Head) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-23

(1) System Installation Annunciator . . . . . . . . . . . . . . . . . . . . . 2-6-24

(2) Remote Tune Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-24

(3) Tuning Cursor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-24

(4) NAV AUDIO On Annunciator . . . . . . . . . . . . . . . . . . . . . . . . 2-6-24

(5) Emergency (EMRG) Mode Annunciator . . . . . . . . . . . . . . . . 2-6-24

(6) Squelch (SQ) Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25

(7) Transmit (TX) Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25

(8) NAV AUDIO On/Off Switch . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25

(9) Squelch (SQ) On/Off Switch . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25

(10) Tuning Knobs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25

(11) Normal/Emergency Mode Switch . . . . . . . . . . . . . . . . . . . . 2-6-25

(12) Transfer Button . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25

(13) Radio Tuning Annunciators . . . . . . . . . . . . . . . . . . . . . . . . 2-6-25

C. AV-850A Audio Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-26

(1) COM1, COM2, COM3, and HF Microphone Switches . . . . . . 2-6-27

(2) Passenger Address (PAX) Microphone Switch . . . . . . . . . . 2-6-27

(3) Emergency (EMER) Switch . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-27

(4) Audio Source Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-27

(5) ID/Voice Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-28

(6) Speaker and Headphone Controls . . . . . . . . . . . . . . . . . . . 2-6-28

(7) Sidetone (ST) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-28

(8) Marker (MKR) Beacon Volume Control . . . . . . . . . . . . . . . . 2-6-28

(9) Marker Beacon MUTE and HI/LO SENS Control . . . . . . . . . . 2-6-28

(10) Interphone (INPH) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-28

D. RCZ-851(x) Integrated Communications Unit . . . . . . . . . . . . . . . . . 2-6-29

E. RNZ-851(x) Integrated Navigation Unit . . . . . . . . . . . . . . . . . . . . . . 2-6-32

F. AT-860 ADF Combined Sense/Loop Antenna . . . . . . . . . . . . . . . . . 2-6-35





3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-36

A. RM-855 Radio Management Unit Interface Diagram . . . . . . . . . . . . 2-6-38

B. CD-850 Clearance Delivery Head Interface . . . . . . . . . . . . . . . . . . . 2-6-43

C. AV-850A Audio Control Unit Interface . . . . . . . . . . . . . . . . . . . . . . 2-6-44

D. Communications Unit Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-53

E. Navigation Unit Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-57



B. RMU Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-61




FIGURE/TITLE PAGE


Figure 2-6-1. Typical RM-855 Radio Management Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-2

Figure 2-6-2. Backup Navigation Display Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-10

Figure 2-6-3. Typical TO Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-11

Figure 2-6-4. Typical FROM Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-12

Figure 2-6-5. Typical ILS Format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-13

Figure 2-6-6. 90-Degree Intercept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-14

Figure 2-6-7. VOR Information Not Available . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-15

Figure 2-6-8. ILS Information Not Available . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-16

Figure 2-6-9. ADF Information Not Available . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-17

Figure 2-6-10. RSB Information Not Available . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-18

Figure 2-6-11. Heading Information Not Available . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-19



Figure 2-6-14. Typical CD-850 Clearance Delivery Control Head . . . . . . . . . . . . . . . . . . 2-6-23

Figure 2-6-15. AV-850A Audio Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-26

Figure 2-6-16. RCZ-851(x) Integrated Communications Unit . . . . . . . . . . . . . . . . . . . . . . 2-6-29

Figure 2-6-17. RNZ-851(x) Integrated Navigation Unit . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-32

Figure 2-6-18. AT-860 ADF Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-35

Figure 2-6-19. Radio System Data Buses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-36

Figure 2-6-20. Radio Management Unit Interface Diagram . . . . . . . . . . . . . . . . . . . . . 2-6-41/42

Figure 2-6-21. CD-850 Clearance Delivery Control Head Interface Diagram . . . . . . . . . . 2-6-43




Figure 2-6-23. Communication Unit Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . 2-6-55/56

Figure 2-6-24. Navigation Unit Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-59/60

Figure 2-6-25. PFD Radio System Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-62

Figure 2-6-26. RMU Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-63



List of Tables

TABLE/TITLE PAGE


Table 2-6-1. RM-855 Radio Management Unit Leading Particulars . . . . . . . . . . . . . . . . . . 2-6-3

Table 2-6-2. CD-850 Clearance Delivery Control Head Leading Particulars . . . . . . . . . . 2-6-23

Table 2-6-3. AV-850A Audio Control Unit Leading Particulars . . . . . . . . . . . . . . . . . . . . 2-6-26

Table 2-6-4. RCZ-851(x) Integrated Communications Unit Leading Particulars . . . . . . . . 2-6-30

Table 2-6-5. RNZ-851(x) Integrated Navigation Unit Leading Particulars . . . . . . . . . . . . 2-6-33

Table 2-6-6. AT-860 ADF Antenna Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . 2-6-35



SECTION 2.6SRZ-850 INTEGRATED RADIO SYSTEM

1. General

The PRIMUS® 1000 Integrated Avionics System installed in the Embraer 145 aircraft includesthe SRZ-850 Integrated Radio System. This radio system is also known as the PRIMUS® IIIntegrated Radio System. It is a dual system, with system No. 1 being the left side and systemNo. 2 being the right side.

The radio system provides communication and navigation information for the flight crew. Thenavigation data is fed to the integrated avionics computers (IACs) for display on the electronicdisplay system (EDS), and for use by the automatic flight control system (AFCS), both of whichare resident within the IACs. Navigation data is also fed to the flight management system(FMS) computer.

The integrated radio system has a number of options which are selected by the installer bymeans of configuration straps. The strap options for the NAV and COM units are programmedon a strap assembly which is aircraft unique and is electrically connected to each unit. Strapoptions for the RMU and CDH are pin programmed on the unit mating connector. There is nopin programming in the audio system. Strap programming procedures are covered inSECTION 3, INTERCONNECTS, of this manual.

The integrated radio system also provides interface with and control of the traffic alert andcollision avoidance system (TCAS II).

The SRZ-850 Integrated Radio System standard installation consists of the following LRUs:

• Two RM-855 Radio Management Units (RMUs)

• One CD-850 Clearance Delivery Control Head (CDH)

• Three AV-850A Audio Control Units

• One RCZ-851E Integrated Communications Unit (VHF COM and DIVERSITY MODE S) -Left side

• One RCZ-851G Integrated Communications Unit (VHF COM and ATCRBS) or oneRCZ-851H Integrated Communications Unit (VHF COM only) - Right side

• One RNZ-851 Integrated Navigation Unit (NAV, DME, and ADF) - Left side

• One RNZ-851C Integrated Navigation Unit (NAV only) - Right side

• One AT-860 ADF Combined Sense/Loop Antenna - Left side.

Available options include the following LRUs:

• One RCZ-851E Integrated Communications Unit (VHF COM and DIVERSITY MODE S) -Right side

• One RNZ-851 Integrated Navigation Unit (for second ADF and second DME) - Right side

• One AT-860 ADF Combined Sense/Loop Antenna (for second ADF) - Right side.




A. RM-855 Radio Management Unit

Figure 2-6-1 shows a graphical view of a typical RM-855 Radio Management Unit (RMU).Table 2-6-1 provides items and specifications that are particular to the unit.

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

COM 1 NAV1

MEMORY-3

ATC/TCAS ADF1

1 TA/RA

RANGE:

123.20131.27

1471

NORMAL

MEMORY-1

110.30

109.35

ADF

162.5

6

CURSOR

FUNCTIONKEYS

TRANSFER("FLIP-FLOP")KEY

LINE SELECTKEY

TUNINGKNOBS

PHOTO SENSOR

AD-51053@

TCAS DSPY 1

Figure 2-6-1. Typical RM-855 Radio Management Unit



Table 2-6-1. RM-855 Radio Management Unit Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.06 in. (128.5 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.06 in. (103.1 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.10 in. (281.9 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.2 lb (3.27 kg)

Power Requirements . . . . . . . . . . . . . . . . . . . . . 28 V dc, 74 W (max)


• Knob, Large . . . . . . . . . . . . . . . . . . . . . . . . HPN 7012149

• Knob, Small . . . . . . . . . . . . . . . . . . . . . . . . HPN 7013773

• Setscrew, 4-40 × 1/8-inch, cup point steel (4) HPN 2500148-128

Mating Connectors . . . . . . . . . . . . . . . . . . . . . . MS3126F24-61S

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Clamp, HPN 7000066-11

The radio management unit (RMU) is the central control unit for the entire radio system. Itprovides complete capability for controlling the operating mode, frequencies, and codeswithin all the units of the radio system. Additionally, the RMU has the capability to switchits operation from its primary radio system to the cross-side system. The RMU is a colorliquid crystal display (LCD) based controller featuring the proven concept of selecting afunction by pushing a line select key adjacent to the parameter that is to be changed. Anyselectable parameter, such as a VOR frequency, may be changed by pressing thecorresponding line key next to the displayed parameter and then rotating the controllertuning knob. For some functions, additional pushes of the line select key will toggle modesor recall stored numbers.

The RMU is also the input to the radio system for external FMS tuning in that digital signalsfrom the FMS come into the RMU where they act in much the same manner as if the fronttuning knob was being operated. This allows the FMS to enter into the system in anorganized manner, and will appear to the system as if the flight crew is tuning the receiver.

During normal operation, complete EICAS is provided to the display system by the DAUsvia ARINC 429. The DAUs also provide backup ARINC 429 inputs to both radiomanagement units (RMUs), for backup display of the essential engine parameters forcontinued safe flight. No other ARINC 429 receivers or 429 transmitters are connected tothese particular data buses. The backup engine parameters are shown on two pages, oneither RMU, in the event of a problem in the EDS. A combination of digital readout of theengine data together with bar graphs depicting rate of change is illustrated in Figures2-6-12 and 2-6-13.

Backup navigation data is provided to the RMU via the radio system bus (RSB) from bothremote navigation units, and/or via RS-422/CSDB data buses from whichever remotenavigation unit is married to the clearance delivery CDH for emergency backup purposes.



The RMUs are located in the center console just forward of the engine power levers.

For ease of operation, the RMU screen is divided into windows. Each window groups thedata associated with a particular function of the radio system. Each window (COM, NAV,ATC, ADF, and TCAS) provides for complete control of frequency and/or operating mode ofthe associated function. The RMU also has other display modes, called pages, whichperform additional features and functions for the control of the radio system. TheATC/TCAS window formats are determined by the actual installation.

Located on the front of the RMU is a button labeled PGE, which, when pressed, causes theRMU to toggle through different pages of the display. The normal five window display iscalled the Main Page and is always present under normal operation. The other pages areassociated with preset memory location and operation for the NAV and the COM windows.Further pages are available by a combination of control buttons and menus to enable thedisplay of various maintenance data from within the radio system. If any page other thanthe Main Page is being displayed, the bottom left line select key is the Return key. In allcases, pressing the return key will change the display back to the Main Page.

Since the TCAS II system is installed, the RMU windows are as shown in Figure 2-6-1. IfTCAS II is not installed, both bottom windows may be disabled. This is done by firstpressing the PGE key, pressing the key adjacent to MAINTENANCE, and pressing the keyadjacent to RMU SETUP. On the RMU SETUP page, pressing the line select key at eitherend of DISPLAY will toggle between ENABLED and DISABLED. Pressing the line selectkey at the lower left will return the display to the main page. This selection function is onlyavailable when weight on wheels is true.

The ATC FLIGHT ID function is designed for use by the airlines. In the Embraer 145, thisfunction should be enabled. In the event that it is inadvertently disabled, the Flight IDdisplay line at the bottom of the transponder window on the main page will be absent.

As with the operation of the cursor in the COM and NAV windows, pressing the keyadjacent to the upper half of the transponder window connects the TUNE knobs to thenumbers. Large knob = left two numbers, small knob = right two numbers.

Pressing the next lower line select key moves the cursor to the bottom half of thetransponder window. This key now has a toggle function between active and standby.When active, the mode of operation is changed by rotation of one of the tune knobs.

With TCAS installed, the modes are:

• ATC ON Replies on Modes S and A, with no altitude reporting

• ATC ALT Replies on Modes A,C and S, with altitude reporting

• TA ONLY The TCAS traffic advisory mode is enabled

• TA/RA The TCAS traffic advisory mode and the resolution advisory mode are bothenabled.



There is a secondary function available when the cursor is in the lower transponderwindow. Pressing the 1/2 key on the lower RMU panel will select which transponder isactive. With TCAS installed, the banner line at the top of the transponder window willindicate ATC/TCAS, and the modes will be those listed above for with TCAS.

The remaining two line select keys move the cursor to either the RANGE line or to thevertical window line. Repeated presses of the RANGE line select key will toggle throughthe range selections (6, 12, 20, 40). The TUNE knob will also change the range when thecursor is in the RANGE window. The vertical window displays are: NORMAL, ABOVE,BELOW, and may be selected by repeated pressing of the line select key, or by rotation ofthe TUNE knob.

Further details of the RMU operation of the TCAS subsystem is described in CHAPTER 8of this section of this manual.

Circuitry within the RMU has been designed to control the light intensity and colors of theRMU and to provide for the ultimate in color tracking across various levels of brightness.The LCD brightness is adjusted by pushing the dim button on the front of the panel andusing the tuning knob in the same manner as other functions are selected. Also providedis a photo sensor on the front of the RMU which senses the ambient light condition andcan adjust the RMU intensity to compensate for varying levels of light as the aircraftmaneuvers in the sunlight. This is a feature which will keep the readability of the RMU at ahigh level while not requiring the pilot to turn the intensity up and down each time the panelpasses from shadow to sunlight.

As a safety feature of the RMU, should any of the components of the radio system fail torespond to commands from the RMU, the frequencies or operating commands associatedwith that particular function will be removed from the RMU and replaced with dashes. Thiswill alert the crew to the fact that the radio system operation is not normal.

Also available in the RMU is a maintenance mode of operation, when not in flight. Duringthis mode, various pages are utilized to allow maintenance personnel access to themaintenance log data and operating conditions of the radio system. In the AircraftMaintenance Mode (AMM), parameters may be examined but may not be modified in anyway by the crew.



(1) RMU Controls

The following paragraphs describe each control on the RMU.

(a) Photo Sensor

The photo sensor senses the ambient light and causes the LCD brightness tobe automatically adjusted to compensate for varying levels of light as theaircraft maneuvers in the sunlight.

(b) Transfer Keys

The transfer key when pushed, flip-flops the active frequency (Top Line) andthe preset frequency (Bottom Line) of the COM or NAV window. Pressing bothtransfer keys simultaneously, while on the ground, allows entry into the aircraftmaintenance mode. For further information about the Aircraft MaintenanceMode, refer to SECTION 7 of this manual.

(c) Line Select Keys

The first push of the line select key moves the yellow cursor to surround thedata field associated with that particular line select key. This thenelectronically connects that data field to the tuning knobs so that frequency ormode may be changed. For some functions, additional pushes of the lineselect key will toggle modes or recall stored frequencies. The line select key,if pressed and held for certain functions, allows ADF and ATC memories to berecalled, and to enter and exit direct tune mode for the COM and NAV.

(d) Tuning Knobs

The tuning knobs are used to modify the data field enclosed by the cursor.This may be frequency or mode depending upon the selected data field.

(e) Squelch (SQ) Key

Pressing the SQ key causes the COM radio to open its squelch and allow anynoise or signal present in the radio to be heard in the audio system. Thesquelch key is strictly a toggle. Pressing the key toggles SQ. The letters SQare annunciated along the top line of the COM window when the squelch isopened by using this key.

(f) Dimming (DIM) Key

Pressing the DIM key connects the RMU brightness control to the tuning knoballowing the display to be adjusted to match overall cockpit brightness.



(g) Cross-Side (1/2) Key

With the cursor in any window except the ATC or TCAS display, pressing thiskey transfers the entire RMU operation and display to the cross-side system.The legend color changes from white to magenta when the RMU is displaying,and is in control of, data associated with the cross-side system. If the cursoris in the ATC or TCAS display window, pressing this key selects whichtransponder will be in operation.

(h) Store (STO) Key

Pressing the STO key causes a temporary (TEMP) COM/NAV preselectfrequency to be stored in memory and assigned a numbered location, providedthe cursor has first been placed around the TEMP display. The ADF and ATCeach have one memory location. Pressing the STO key causes the currentADF frequency or ATC code to be stored in memory, provided the cursor hasfirst been placed around that frequency or code.

(i) Identification (ID) Key

Pressing the ID key places the transponder in the identification responsemode. The ID squawk terminates after 18 seconds. The identificationresponse mode can also be invoked by the control yoke mounted pushbuttons.

(j) Page (PGE) Key

Pressing the PGE key once will change the RMU display to the RMU PageMenu, except when operating in the Aircraft Maintenance Mode (AMM).Pressing the PGE key a second time will have no effect. When not on theMain Operating page, the RMU assigns a Return function to the lower left lineselect key. Pressing this key will return to the Main Operating page.

(k) Test (TST) Key

Pressing the TST key causes the component associated with the yellowcursor’s present position to activate its internal self-test circuits for a completeend-to-end test of the function. Hold the TST key down for the duration of thetest, about two seconds for COM transceiver, five to seven seconds for DME,ATC, ADF, and about 20 seconds for NAV (VOR/ILS). Releasing the TST keyat any time immediately returns the function to normal operation. The TCAStest function is slightly different, and is described in Chapter 8, SECTION 2 ofthis manual.

(l) DME Key

The DME key deslaves the DME from the active VOR frequency, and allowstuning of a different DME channel without changing active VOR. Successivepresses of the DME key enable display and selection of the DME channels inVHF and TACAN formats.



(2) Backup Navigation Display

The following paragraphs describe the Backup Navigation display capability ofthe RMU. (See Figures 2-6-2 thru 2-6-11.)

(a) The backup navigation display page is selected by the flight crew. Pressingthe PGE key on the RMU brings up the RMU Page Menu. Pressing the lineselect key next to the NAVIGATION legend will cause the backup navigationpage to be displayed.

(b) An overview of the backup navigation page is shown in Figure 2-6-2. Allpossible annunciation fields are shown (which is not representative of anactual display scenario).

The following information is displayed when appropriate:

1 NAV

Operating frequency of the No. 1 NAV receiver. Pressing the line selectkey next to NAV moves the yellow cursor to surround the NAV frequency.The TUNE knobs now act to control the NAV frequency. If the frequency ischanged by any other source, the change will be echoed on the backupnavigation display, now in yellow to indicate that another source did thetuning.

The yellow cursor box returns to the CRS window 20 seconds after the lasttuning command is entered.

2 ADF

Operating frequency of the No. 1 ADF receiver. Pressing the line selectkey next to ADF moves the yellow cursor to surround the ADF frequency.The TUNE knobs now act to control the ADF frequency. If the frequency ischanged by any other source, the change will be echoed on the backupnavigation display, now in yellow to indicate that another source did thetuning.

3 Bearing pointers for VOR or ADF (or single and double arrows whenpointers are out of view)

4 Digital ADF and VOR bearing readouts and their "diamond" and "circle"identifiers



5 CRS

Selected Course. Pressing the line select key next to CRS moves theyellow cursor to surround the CRS readout (if it is not already there). TheTUNE knobs now act to select the desired course. The large knobchanges the course in 10-degree increments, and the small knob changesthe course in 1-degree increments.

The course pointer moves as the course selection changes.

6 TO/FROM indication

7 DME

DME distance to tuned station.

8 Marker beacons

9 Magnetic heading from No. 1 AHRS via the DAUs (ARINC 429)

10 Lateral deviation (VOR and Localizer)

11 Vertical deviation (Glideslope only)

In the event that RSB information is unavailable, VOR/ILS only is receivedby the RMUs via the CSDB data line directly from the No. 1 NAV receivermodule. Magnetic Heading is ARINC 429, and CRS is generated withinthe RMU, therefore these two displays are not affected by a loss ofRSB data.



333N

DIGITALHEADING

ADFFREQUENCY(WHITEOR YELLOW)

VORPOINTER"TAIL"AND "HEAD"(GREEN)

C R S

T O 07.2

429 .5112.70

3 3 3

ADF

VOR

195˚

065˚

NAV ADF

DME

O M3 60

3 60SELECT-TO-TUNE BOX(YELLOW)

VOR POINTER OUTOF VIEW LEFT (CYAN)

ADF POINTER TOLEFT (WHITE)

COURSE POINTER(GREEN)

DIGITAL VORBEARING (CYAN)

DIGITAL ADFBEARING (WHITE)

LATERAL DEVIATIONINDICATION (WHITE)

SELECTED COURSE(GREEN)

VOR POINTER OUTOF VIEW RIGHT(CYAN)

ADF POINTER TORIGHT (WHITE)

HEADING COMPASS(WHITE)

GLIDESLOPEPOINTER (GREEN)

GLIDESLOPESCALE (WHITE)

LATERALDEVIATION SCALE(WHITE)

DME DISTANCE(GREEN)

AIRCRAFT SYMBOL(WHITE)

TO/FROM INDICATOR(GREEN)

MARKERS(CYAN, YELLOW, WHITE)0M MM IM

ADF POINTER"HEAD"AND"TAIL"(WHITE)

NAVFREQUENCY(WHITEOR YELLOW)

AD-33602-R1@

T RU

"TRUE" HEADINGINDICATOR (YELLOW)

Figure 2-6-2. Backup Navigation Display Overview



A typical VOR navigation format is shown in Figure 2-6-3. The "head" ofthe VOR Bearing pointer is a 005 degrees and the aircraft heading andcourse selection is at 000 degrees. Therefore, the course deviation of 5degrees is shown by the course deviation bar to be 1 dot to the right of theaircraft symbol. TO is annunciated at the bottom of the display.

The double arrow in the upper left part of the display indicates that theADF pointer is to the left of the displayed 90-degree sector.

For DME distances greater than 99.9 nautical miles, the distance isdisplayed to the nearest whole number. Figure 2-6-3 shows a DMEdistance of 150 nautical miles.

AD-33593@

429.5109.90 NAV ADF

360

CRSTO

N33 3

DME 150360

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

VOR 005

300 ADF

Figure 2-6-3. Typical TO Format



Figure 2-6-4 shows a typical FROM format. The selected course is 015degrees and the "tail" of the VOR Bearing pointer is at 005 degrees.Therefore, there are 2 dots of course deviation displayed. FR isannunciated at the bottom portion of the display.

The "tail" of the ADF pointer is shown at 340 degrees, and the DMEdistance is 7.5 nautical miles.

AD-33594@

CRSFR

429.5112.70

N33 3

NAV ADF

ADF

VOR

DME

185˚

160˚

07.5

360

015

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

Figure 2-6-4. Typical FROM Format



Figure 2-6-5 shows a typical ILS format. Notice that all VOR indicationsare absent.

MM in the bottom portion of the display indicates that the Middle Marker isbeing received.

The aircraft is tracking the glideslope and is slightly to the right of thelocalizer beam.

The "head" of the ADF Bearing pointer is at 330 degrees.

AD-33595@

CRSMM

429.5109.90

N33 3

NAV ADF

ADF

DME

330˚

7.2

360

360

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

Figure 2-6-5. Typical ILS Format



Figure 2-6-6 shows a typical 90 degrees intercept of the localizer frombelow the glideslope. The glideslope pointer and the localizer deviationbar are pegged indicating deviations greater than the display allows.

The "head" of the ADF Bearing pointer is at 000 degrees.

AD-33596@

CRS

429.5109.90

N33 3

NAV ADF

ADF

DME

000˚

25.2

360

090

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

Figure 2-6-6. 90-Degree Intercept



Figure 2-6-7 shows the display when VOR information is not available, oneither the RSB or CSDB.

The digital bearing for the VOR is removed.

The VOR Bearing pointer is removed.

The lateral deviation bar is removed and the scale is overwritten with a redX.

The DME distance is replaced with yellow dashes.

The TO/FROM indication is removed.

AD-33597@

429.5112.70 NAV ADF

360

CRS

N33 3

ADF

VOR

DME - - -360

300˚

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

Figure 2-6-7. VOR Information Not Available



Figure 2-6-8 shows the display when ILS information is not available, oneither the RSB or CSDB.

The lateral deviation bar is removed and the scale is overwritten with ared X.

The glideslope pointer is removed and the scale is overwritten with ared X.


AD-33598@

CRS

429.5109.90

N

33 3

NAV ADF

ADF

DME

330˚

- - -

360

360

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

Figure 2-6-8. ILS Information Not Available



Figure 2-6-9 shows the display when ADF information is not available.

The digital bearing for ADF is removed.

The ADF Bearing pointer is removed.

AD-33599@

CRSMM

429.5109.90

N

33 3

NAV ADF

ADF

DME 7.2

360

360

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

Figure 2-6-9. ADF Information Not Available



Figure 2-6-10 shows the display when RSB information is not available.

The ADF frequency is replaced by white dashes.

The ADF digital bearing and the ADF Bearing pointer are removed.


All VOR/ILS information is still available via CSDB from No. 1 NAV receivermodule.

AD-33600@

CRSMM

109.90

N33 3

NAV ADF

ADF

DME

360

360

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

Figure 2-6-10. RSB Information Not Available



Figure 2-6-11 shows the display when heading information is not available.

AD-33601@

CRSMM

429.5109.90

N33 3

NAV ADF

ADF

DME 3.1

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

HDGFAIL

Figure 2-6-11. Heading Information Not Available



(3) Backup Engine Display

The following paragraphs describe the Backup Engine display capability of the RMU.(See Figures 2-6-12 and 2-6-13.)

(a) The backup engine display page may be selected by the flight crew on eitherRMU. Pressing the PGE key on the RMU brings up the RMU Page Menu.Pressing the line select key next to the ENGINE PG1 or ENGINE PG2 legendwill cause the selected backup engine pages to be displayed.

Moving between page 1 and page 2 is done by pressing the lower left lineselect key.

(b) The backup engine display page 1 will be displayed automatically, on the leftside (No. 1) RMU, if both IACs determine that the primary EICAS display is notavailable.

(c) Once the backup engine display has been selected (either automatically or bythe flight crew), the only way to deselect the display is by pressing the PGEfunction key. Additionally, if page 2 has been selected, after 20 seconds thedisplay will automatically return to page 1.

(d) When the backup engine display has been selected automatically, removal issubject to different conditions. The RMU is allowed to page away from enginepages. However, if the backup display enable discretes are active, the RMUwill return to the engine page No. 1 twenty seconds after the last manual inputwas made to the RMU.

If both of the display enable discretes change back to a "primary EICASavailable" condition (for greater than two seconds), the No. 1 RMU changesback to the Main Tuning Page automatically.



Figure 2-6-12 shows backup engine page No. 1.

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

AD-51054@

MAX-TO MAX-TO

102.5 99.9N1

ITTN2

52095.0

850

250

145

PAGE 2 4 MSGS

145

195

910

96.7490

FF PPH

OIL P

OIL T

THRUST RATINGMODE:MAX-TOALT-TO, FLX-TO,CON, CLB, CRZ(CYAN)

LABELS/LEGENDS(WHITE)

N1 POINTER ANDDIGITS (GREEN,RED)

ITT DIGITS(GREEN, RED)

N2 DIGITS(GREEN, RED)

FUEL FLOWDIGITS(GREEN)

OIL PRESSUREDIGITS (RED,AMBER, GREEN)

OIL TEMP DIGITS(RED, AMBER,GREEN)

PAGE 2SELECTION(GREEN)

NUMBER OFCAS MESSAGESON PAGE 2 (WHITE)

Figure 2-6-12. Backup Engine Page No. 1

The following parameters are displayed for both engines.

• N1: Engine Fan Speed

• ITT: Inter Turbine Temperature

• N2: Engine Turbine Speed

• FF PPH: Fuel Flow in Pounds Per Hour

• OIL P: Oil Pressure

• OIL T: Oil Temperature.

The pointers on the engine N1 bar graph displays change color to red whenengine fan speed reaches 100%, and change back to green when engine fanspeed falls below 99.9%. The pointers are cut to ½ triangle at the top of thescale.



When the MSG display flashes in yellow it indicates that there is a newmessage on page 2.

The number to the left of the MSG legend indicates how many messages areon page 2.

Figure 2-6-13 shows backup engine page No. 2.

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

AD-51055@

1200

PAGE 1

MESSAGE LINE #7

LEGENDS/LEGENDS(WHITE)

FUEL QUANITY DIGITS(RED, AMBER, GREEN)

FLAP POSITION DIGITS(GREEN)

PAGESELECTION(GREEN)

MESSAGE LINE #6MESSAGE LINE #5MESSAGE LINE #4MESSAGE LINE #3MESSAGE LINE #2

LG DOWN LOCKED

FQ LBFLAPS

2100

22

CASMESSAGES(GREEN)

Figure 2-6-13. Backup Engine Page No. 2

The following parameters are displayed:

• FQ LB: Fuel Quantity in Pounds

• FLAPS: Flap Position in Degrees.

The seven lines that follow the above information on page 2 are used todisplay up to seven CAS messages as appropriate.



B. CD-850 Clearance Delivery Control Head (Tuning Backup Control Head)

See Figure 2-6-14 and Table 2-6-2.

AD-46688@

SQ NAVAUDIO

MODES

COMNAVNAV AUDIO TX SQ PRENAV

RMT EMRG

123EMERGENCYMODEANNUNCIATOR

SQUELCHANNUNCIATOR

TRANSMITANNUNCIATOR

NAV AUDIOON/OFF SWITCH

SQUELCHON/OFF SWITCH

TUNING KNOBS

SYSTEM INSTALLATIONANNUNCIATOR

REMOTE TUNEANNUNCIATOR

TUNINGCURSOR

NAV AUDIO ONANNUNCIATOR

RADIO TUNINGANNUNCIATORS

TRANSFER KEY

NORM/EMERGENCYMODE SWITCH

Figure 2-6-14. Typical CD-850 Clearance Delivery Control Head

Table 2-6-2. CD-850 Clearance Delivery Control Head Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . 2.62 in. (66.7 mm)

• Width . . . . . . . . . . . . . . . . . 2.38 in. (60.3 mm)

• Length . . . . . . . . . . . . . . . . 7.54 in. (191.5 mm)

Weight . . . . . . . . . . . . . . . . . . . . 1.25 lb (0.567 kg)



• Knob, Mode . . . . . . . . . . . . . HPN 800B0718

• Knob, Coarse Tuning . . . . . . HPN 800B0714

• Knob, Fine Tuning . . . . . . . . HPN 800B0715

• Setscrew, 4-40 × 3/32-inch (6) HPN 100A4634-01

Mating Connectors . . . . . . . . . . . MS3126F20-41SW, HPN 4000809-607

Mounting . . . . . . . . . . . . . . . . . . Panel Mount



The CD-850 Clearance Delivery Control Head (CDH) provides an alternate or emergencybackup capability for tuning the No. 2 VHF COM Module and the No. 2 VHF NAV ReceiverModule, on private line data buses which remain operational, in the event that the primaryRadio System Bus (RSB) tuning is not available, or if the pilot/operator wishes to overridethe bus tuning for any reason. The CDH listens on the RSB and displays the activefrequencies of these two modules.

The CDH is located on the center console in the cockpit of the aircraft, between the RMUs,just forward of the engine power levers.

The CDH uses a transflective, dichroic (black dye), liquid crystal display to provideenhanced readability and reliability. The panel lettering and buttons are internally lighted,using aviation blue-white lighting.

The CDH has several strap options which are used to select various features of the COMor VOR radios in the emergency mode. These features are set at the rear connector viajumper straps. In the Embraer 145, the CD-850 operating mode, as selected by installationstrapping on the rear connector, is the Clearance Delivery mode.

The normal and emergency modes are submodes that are selected by the mode knob.

The following paragraphs describe each function or control:

(1) System Installation Annunciator

The 2 annunciator is ON to indicate that the CD-850 is connected to the No. 2 COMand No. 2 NAV.

(2) Remote Tune Annunciator

This annunciator is inactive in the Embraer 145.

(3) Tuning Cursor

The tuning cursor annunciator is a lighted triangle which is controlled by the transferkey. It indicates which frequency can be changed by the tuning knobs.

(4) NAV AUDIO On Annunciator

This annunciator indicates when NAV audio has been selected on. See paragraph2.C.(3) in this chapter for additional details regarding the emergency audio systemselection and operation.

(5) Emergency (EMRG) Mode Annunciator

This annunciator indicates when the CD-850 has been placed in the emergencyback-up mode which locks out all other COM and NAV tuning sources for the No. 2COM and No. 2 NAV. The No. 2 COM and the No. 2 NAV are now tuned exclusivelyby the CD-850. This annunciator is not related to the emergency frequency of121.5 MHz.



(6) Squelch (SQ) Annunciator

This annunciator indicates that the squelch is opened by the SQ On/Off switch.

(7) Transmit (TX) Annunciator

This annunciator indicates when the COM transmitter is ON.

(8) NAV AUDIO On/Off Switch

This alternate action button is used to toggle NAV audio ON or OFF.

(9) Squelch (SQ) On/Off Switch

This alternate action button is used to toggle the COM squelch ON or OFF.

(10) Tuning Knobs

The tuning knobs are used to change the frequency indicated by the tuning cursor.Large knob = left two numbers, small knob = right two numbers.

(11) Normal/Emergency Mode Switch

This rotary switch knob provides alternate selection of the Normal andEmergency modes.

(12) Transfer Button

The transfer button alternately selects either the COM frequency (top) or the NAVfrequency (bottom) to be connected to the tuning knobs.

(13) Radio Tuning Annunciators

These two annunciators (COM, NAV) are annunciated individually, together with thetuning cursor, to identify the frequency at the top and bottom lines.



C. AV-850A Audio Control Unit


HFCOM3

STSPKR HDPH

DME1 DME2 ID/VOICE

COM1 COM2 PAX

MICROPHONE EMER

ADF1NAV2

MKR

NAV1

MUTE

ADF2

INPH

AD-50461@

Figure 2-6-15. AV-850A Audio Control Unit

Table 2-6-3. AV-850A Audio Control Unit Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . 3.00 in. (76.2 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . 5.75 in. (146.1 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . 7.10 in. (180.3 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.22 lb (1.50 kg)

Power Requirements . . . . . . . . . . . . . . . . . . 28 V dc, 28 W (max)


• Knob, Speaker or Headphone . . . . . . . . HPN 7511039

• Setscrew, 2-56 × 7/8-inch, cup point steel HPN 2500148-64

Mating Connectors:

• J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS27473E20-A41S, HPN 2500981-195

• J2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS27473E20-A41SB, HPN 2500981-197

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . Standard Dzus Rail



The AV-850A Audio Control Unit (Figure 2-6-15) receives digitized audio from remote radiounits through two high speed digital audio buses. The audio control unit decodes thedigital data, controls the gain (volume) of the various channels, adds the channels together,does various filter functions on the audio, and outputs the audio to a digital-to-analogconverter. It contains hardware for switching microphones to various radios, and hardwarefor the interphones as well as the passenger cabin audio and intercoms. Amplifiers arealso provided for driving headphones and speakers.

Two audio control units are located on the aircraft instrument panel, outboard of the EDSdisplays. The third audio control unit is located at the observer’s station.

The following paragraphs describe the audio control unit switch and control functions.

(1) COM1, COM2, COM3, and HF Microphone Switches

These switches when pressed automatically select the desired microphone and atthe same time enable the receiver audio associated with that microphone, regardlessof the setting of the COM or HF audio on/off controls that are located under theswitch.

(2) Passenger Address (PAX) Microphone Switch

When the PAX switch is pressed, the microphone is connected directly to the PAamplifier independent of power being applied to the audio [panel. Microphone biasvoltage is derived from the PA amplifier. Passenger Address sidetone is internallygenerated within the audio panel and is controlled by the INTERNAL SIDETONEside-panel potentiometer and the HEADPHONE volume control while the speakersidetone audio is controlled by the INTERNAL SIDETONE potentiometer, theSPEAKER SIDETONE volume control and the SPEAKER volume control. Duringaudio panel power loss, no PA sidetone will be available. All other audios aredeselected during passenger address operation except for the warning audios.

(3) Emergency (EMER) Switch

When the EMER switch is pressed, the microphone is connected directly to VHFCOM transceiver No. 2, and the transceivers received audio is connected directly tothe aircraft’s headphone. The No. 2 VOR/ILS audio is also connected directly to theaircraft’s headphone if it has been selected by the NAV AUDIO pushbutton on theCDH. When EMER is selected, headphone volume is controlled by the masterheadphone volume control. This mode also disables all other audio control unitmodes.

(4) Audio Source Control

Each control (COM, HF, NAV, ADF, DME) combines the function of switch andvolume control.

Each source control (e.g., COM) energizes a particular channel’s audio whenunlatched (out position) and de-energizes the audio when latched (in position).Rotation of this control will adjust the audio level from minimum at the fully CCWposition to maximum at the fully CW position.



(5) ID/Voice Control

The ID/Voice control is used to filter the VOR and ADF audio signals. In the IDposition (latched in), the VOR or ADF audio is filtered in such a way as to enhancethe Morse Code identification. In the VOICE position (latched out), the audio isfiltered to reduce the Morse Code signal for received VOR/ILS audio. ADF audiopasses through without change when VOICE mode is selected.

(6) Speaker and Headphone Controls

These controls are used to adjust the overall speaker and headphone volume. Theywork in series with the individual audio source controls.

(7) Sidetone (ST)

The sidetone (ST) level control is used to prevent undesirable feedback of speakersidetone audio into the transmitting microphone. Whenever transmitting, both theon-side and off-side speaker sidetone audio levels will be lowered. In addition, theoff-side headphone and sidetone levels are lowered. When the ST control is latched(in position) the cockpit speaker is turned OFF. When the control is unlatched (outposition) the cockpit speaker is ON.

(8) Marker (MKR) Beacon Volume Control

The marker volume control is a latched switch used to control the marker audiovolume. It differs from the other volume controls in that it may not be turned downbelow a level which is adjusted by a potentiometer inside the audio control unit. Thisprevents the marker audio from being turned down too low to be heard, causing themarker signal to be missed.

(9) Marker Beacon MUTE and HI/LO SENS Control

The marker beacon receiver sensitivity is controlled by rotating the control. Pressingthe control activates the marker mute function that is used to temporarily silence theMarker Beacon audio (non-latching). When the marker audio is muted, it will remainmuted as long as the audio level is above a threshold setting. When the audio leveldrops below the threshold, a time-out sequence begins that continues to mute themarker audio for a fixed period of time. After the time-out the marker audio isunmuted.

(10) Interphone (INPH)

The interphone volume control adjusts the headset audio level when the interphonefunction is used. Normally, interphone audio is available only over the headset. Theinterphone function ties together all three audio control units and any externallylocated maintenance audio jacks.



D. RCZ-851(x) Integrated Communications Unit


AD-13743-R2@

Figure 2-6-16. RCZ-851(x) Integrated Communications Unit



Table 2-6-4. RCZ-851(x) Integrated Communications Unit Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.38 in. (85.9 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.90 in. (226.1 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.01 in. (355.9 mm)

Weight:

• RCZ-851E/G . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.3 lb (5.13 kg)

• RCZ-851H . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TBD

Power Requirements:

• RCZ-851E/G . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 V dc, 227 W (max)

• RCZ-851H . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 V dc, 171 W (max)


• XC-850 Cluster Module (RCZ-851E/G/H) . . . . . . . . . HPN 7510784-904

• TR-850 Comm Module (RCZ-851E/G/H) . . . . . . . . . . HPN 7510764-902

• XS-852 Diversity Transponder Module (RCZ-851E) . . HPN 7517400-902

• XS-850A ATCRBS Transponder Module (RCZ-851G) HPN 7517400-904

Mating Connectors (J1) . . . . . . . . . . . . . . . . . . . . . . . . . HPN 7500294-106

Mounting:

• RCZ-851E/G/H . . . . . . . . . . . . . . . . . . . . . . . . . . . . MT-851 Tray, HPN 7510124-920

• Fan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HPN 7500524-002

• Strap Board Assembly . . . . . . . . . . . . . . . . . . . . . . HPN 7510280-901



The RCZ-851(x) Integrated Communication Unit is a complete self-containedcommunication system. It contains the VHF communication transceiver and the air trafficcontrol transponder. Also, within the communication unit is a cluster module whichcontains the circuitry necessary to handle all of the digital outputs of the communicationmodules and place them on the digital audio and radio system buses. Each one of themodules is self-contained within its own housing, has its own internal power supply andconnects to the cluster module via ribbon cable. Cooling is provided by a noncritical rackmounted fan. Temperature sensors inside the individual modules report temperature riseto the cluster module, which in turn switches the fan on, and monitors its operation. Whenthe temperature drops sufficiently, the fan is switched off.

The COM unit is located in the avionics nose bay.

In the RCZ-851E COM Unit, the cluster module is a single printed circuit board that isattached to the rear connector nearest to the outside of the rack (J1), and the transponderis the Diversity Mode S.

In the RCZ-851G COM Unit, the cluster module connector is the same as in the RCZ-851Eand the transponder is the ATCRBS.

In the RCZ-851H COM Unit, the cluster module connector is the same as in the RCZ-851Eand there is no transponder.

A heat sink is associated with the COM unit and is attached to the front of each of themodules to provide a heat path from the internal structure of the box to the front surfacewhere there is adequate radiating surface provided in the free area. At the rear of thecommunication unit are flush mounted antenna connectors and the aircraft harnessconnector.

The cluster module has its own on-board power supply and receives its primary 28 voltinput power from both the VHF COM transceiver module and the transponder module sothat in the event either of them is energized, the cluster module will be energized. TheCOM cluster module contains audio interface circuitry for the signals from the COM unit.Because of the nature of its operation, the transponder has no audio output circuitry.



E. RNZ-851(x) Integrated Navigation Unit


Figure 2-6-17. RNZ-851(x) Integrated Navigation Unit



Table 2-6-5. RNZ-851(x) Integrated Navigation Unit Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.38 in. (85.9 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.90 in. (226.1 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.01 in. (355.9 mm)

Weight:

• RNZ-851 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.6 lb (6.16 kg)

• RNZ-851C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.4 lb (4.26 kg)

Power Requirements:

• RNZ-851 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 V dc, 69 W (max)

• RNZ-851C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 V dc, 29 W (max)


• XN-850 Cluster Module (RNZ-851/851C) . . . . . . . . . HPN 7510164-921

• NV-850 Nav Module (RNZ-851/851C) . . . . . . . . . . . HPN 7510134-831

• DM-850 DME Module (RNZ-851) . . . . . . . . . . . . . . . HPN 7510184-902

• AD-850 ADF Module (RNZ-851) . . . . . . . . . . . . . . . HPN 7510114-811

Mating Connector . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HPN 7500759-001

Mounting:

• Tray . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MT-851 Tray, HPN 7510124-910

• Fan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HPN 7510295-901

• Strap Board Assembly . . . . . . . . . . . . . . . . . . . . . . HPN 7510280-901



The RNZ-851 Integrated Navigation Unit is a complete self-contained navigation system. Itcontains the NV-850 VHF NAV Receiver, the DM-850 DME, and the DF-850 ADF modules.Also within the RNZ-851 is an XN-850 Cluster Module that provides the interface with theNV-850, DM-850, DF-850 and other units of the integrated radio system, and digitizes thereceived audio for the digital audio system. Cooling is provided by a non-critical rackmounted fan. Temperature sensors inside the individual modules report temperature riseto the cluster module, which in turn switches the fan on, and monitors its operation. Whenthe temperature drops sufficiently, the fan is switched off.

The NAV unit is located in the avionics nose bay.

The RNZ-851C is nearly identical to the RNZ-851, but does not contain ADF or DME.

A heat sink is associated with the NAV unit and is attached to the front of each of themodules to provide a heat path from the internal structure of the box to the front surfacewhere there is adequate radiating surface provided in the free area. At the rear of thecommunication unit are flush mounted antenna connectors and the aircraft harnessconnector.

The cluster module has its own on-board power supply and receives its primary 28 voltinput power from either the VHF NAV receiver module, the DME module or the ADFmodule so that in the event any one of them is energized, the cluster module will beenergized. The NAV cluster module contains audio interface circuitry for the signals fromthe VHF NAV and ADF modules. The morse decoder within the DME module sends onedata bit during morse tones. The cluster module includes this data bit in the digital audiodata, and the tone is recreated by the audio control unit.



F. AT-860 ADF Combined Sense/Loop Antenna


AD-14150@

Figure 2-6-18. AT-860 ADF Antenna

Table 2-6-6. AT-860 ADF Antenna Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . 1.51 in. (38.3 mm)

• Width . . . . . . . . . . . . . 8.33 in. (211.6 mm)

• Length . . . . . . . . . . . . 16.30 in. (414.8 mm)

Weight . . . . . . . . . . . . . . . . 3.7 lb (1.68 kg)

Power Requirements . . . . . . +15 V dc from ADF Receiver

User Replaceable Parts . . . . None

Mating Connectors . . . . . . . Cannon KPT08P12-10S, HPN 7500489-524

Mounting . . . . . . . . . . . . . . Outer Fuselage

Gasket . . . . . . . . . . . . . . . . HPN 7020801-932

The AT-860 ADF Antenna performs the function of reception, amplification, andcombination of RF signals so as to yield low frequency reception and directionalinformation. The antenna also contains a self-test circuit which radiates a 120 kHz signalinto the sense and loop antennas. This checks the operation of both the AT-860 ADFAntenna and the DF-850 ADF Receiver Module. Proper operation is indicated by a 1 kHztone and a bearing indication of 135 degrees relative to the nose of the aircraft.

The ADF antenna is located on the center fuselage on top of the aircraft.



3. Operation

Figure 2-6-19 shows the integrated radio system buses. Command, control, and datacommunications between LRUs is via RSB. RCB within the COM and NAV units is on a ribboncable. Digital audio from the NAV and COM units to the audio control units is via digital audiobuses. Command and control data from the CDH to COM No. 2 and NAV No. 2 is via privateline data buses. All of these buses are described in SECTION 1, SYSTEM OVERVIEW, of thismanual.

RIG

HT

SID

ES

EC

ON

DA

RY

RS

B

LEFT

SID

ES

EC

ON

DA

RY

RS

B

AUDIO BUS NO. 1

AUDIO BUS NO. 2

PRIMARY RSB

AUDIOCONTROL

UNITNO. 2

AUDIOCONTROL

UNITNO. 1

RMUNO. 1

RMUNO. 2

COMMUNITNO. 1

COMMUNITNO. 2

NAVUNITNO. 1

NAVUNITNO. 2

TO IC-600 IACS

ARINC 429

TO/FROM TCAS

AD-50887-R1@

AUDIOCONTROL

UNITNO. 3

ARINC 429

TO/FROM TCAS

CDH

Figure 2-6-19. Radio System Data Buses



Each remote radio unit contains a number of functional modules. These are packagedas follows:

• RCZ-851E Integrated Communications Unit

– TR-850 VHF COM Transceiver Module

– XS-852A Diversity Mode S Transponder Module

– XC-850 COM Cluster Module (RSB and Digitized Audio Interface)

• RCZ-851G Integrated Communications Unit


– XS-850A ATCRBS Transponder Module


• RCZ-851H Integrated Communications Unit



• RNZ-851 Integrated Navigation Unit

– NV-850 VHF NAV Receiver Module

– DM-850 DME Transceiver Module

– DF-850 ADF Receiver Module

– XN-850 NAV Cluster Module (RSB and Digitized Audio Interface)

• RNZ-851C Integrated Navigation Unit

– NV-850 VHF NAV Receiver Module

– XN-850 NAV Cluster Module (RSB and Digitized Audio Interface).

Controls and the associated displays for the radios are available in the following units:

• AV-850A Audio Control Unit

• RM-855 Radio Management Unit (RMU)

• CD-850 Clearance Delivery Control Head (CDH).

The basic radio control functions are:

• VHF COM Mode and Frequency

• VHF NAV Mode and Frequency

• ADF Mode and Frequency

• Transponder Reply Code and Mode

• TCAS Mode, Range and Vertical Window

• DME (Independent Channeling in the Hold Mode)

• Audio Control Unit.



Frequency and mode control of the radios can be input by the operator from either the RMU, theFMS, or the CDH. The CDH is somewhat limited in that it is only connected to COM No. 2 andNAV No. 2. Microphone selection, radio headset and speaker audio selection, and volumecontrol are provided by the audio control unit. Audio switching control is input by means ofcontrols on the audio control unit itself. The received audio signals are transmitted from theremote units to the audio control unit via a dedicated digital audio bus. The microphone audiooutput from the audio control unit to the remote mounted transmitters is analog.

Basic to the overall system design are cluster modules in the COM and NAV remote units. Thecluster module is an interfacing element which collects data from the RSB, distributes this datato the respective functional modules (ADF, DME, etc.) via the radio communication bus (RCB),and also collects data via RCB from the functional modules to be broadcast on the RSB. Thecluster module is also responsible for digitizing the received audio and transmitting the digitizeddata on the digital audio bus.

A. RM-855 Radio Management Unit Interface Diagram

See Figure 2-6-20.

The RMUs broadcast messages addressed to radio functional modules and receive datafrom the radios via the RSB. Three major functions of the RMU are to: output tuning(channel or frequency) control data, output operational mode control data for the radios,and display the tuned active channel or frequency and operational mode.

The marker beacon receiver sensitivity control is a discrete which goes to ground wheneither crew member selects HI sense on the audio control unit. The ground is applied toboth RMUs, which in turn send the appropriate data command out on the RSB, and thatcrew members marker beacon receiver is then switched to the HI sense mode of operation.

ATC Ident input is at ground when a crew member presses either of the yoke mountedremote ident buttons. The ground is applied to both RMUs, which in turn, send theappropriate data command out on the RSB.

Other inputs include power, lighting and ground connections.

Weight On Wheels is at ground when the aircraft is on the ground (WOW is true).

The left DAU supplies information pertaining to the left engine, and the right DAU suppliesinformation pertaining to the right engine. Display format, titles, warning messages, etc.are supplied by BOTH DAUs. This backup function can be enabled by a discrete going toground from either IAC that detects a failure in the EICAS display. It can also be selectedby crew members through the RMU menu pages.

In the event of a total EFIS failure, crew members can select the backup NAV function, andthe RMU will then display the backup NAV function. Although this function is notautomatic, as is the Engine Backup display, it, too, is selected through the RMUmenu pages.

When the optional FMS is installed, command data from the FMS computer is entered intothe radio system through the RMU. This input may be disabled by the flight crew, throughthe RMU menu pages, if required.



On the ground, the RMU will self-initiate a system Power-On Self-Test (POST) when poweris first applied to the radio system and at other times with weight-on-wheels when powerhas been off for more than 10 seconds.

The first page to appear on the RMU screen will be POST in progress. POST lasts45 seconds. During POST, the following may be observed on the horizontal situationindicator display on both primary flight displays (PFDs):

• Marker indicators and tones will annunciate in the order of 3000 Hz (WHITE IM),1300 Hz (AMBER MM), and 400 Hz (BLUE OM).

• Localizer and glideslope deviation bars will indicated centered course for approximately2 seconds with flags out of view

• Localizer and glideslope deviation bars will deflect left (localizer) and up (glideslope)one dot deflection, for approximately 2 seconds with flags out of view

• If the course selector is on zero degrees, VOR deviation bar centers on a course ofzero degrees, TO, and RMI indicates zero degrees north for approximately 5 secondswith flags out of view

• DME TEST will appear

• 10.0 NM, 120 KT, and 5 minutes TTG

• RMI ADF pointer will slew to 135 ± 10 degrees, relative to aircraft heading

• Audio tone is heard through the audio system.

The system also includes a Pilot-Activated Self-Test (PAST), which is accomplished bypressing a line select key to place the cursor in the window for the module to be tested,and pressing and holding the TST button. Details of the PAST are described inSECTION 7, MAINTENANCE TEST AND FAULT ISOLATION, of this manual. For fullmaintenance information on the radio system, refer to Honeywell Pub. No. A15-3800-01,PRIMUS® II Integrated Radio System Operation and Installation Manual.

For full operational information on the RMU, CDH, and audio control units, refer toHoneywell Pub. No. 28-1146-50, PRIMUS® II Integrated Radio System, Pilot’s Manual.



(Blank Page)



ENGINE BU DISPLAY ENABLE NO. 2

AA

t

121J1B144J1 144J1

(H)

(L)

RM-855 RADIO MANAGEMENTUNIT NO. 1

NZ-2000 NAVIGATIONCOMPUTER

28 V DC AVIONIC ESSENTIAL BUS

28 V DC BUS 1

DC GROUND

DIMMING BUS

LIGHTING GROUND

C

LL

D

MM

G

H

+5 V LIGHTING

LIGHTING RETURN

40

41

136J1B

(H)

(L)

DA-800 DAU UNITNO. 1

CC

DD

FF

GG

L

160J2

MKR SENSE

hi

MKR SENSE LEFTMKR SENSE RIGHT

V

W

M

N

S

P

R

U

Y

b

c

d

wx

ATC IDENT

W.O.W.

W.O.W. POLARITY

SIGNAL GROUND

ON/OFF PAGE DISABLE

SIDE SELECT B0

SIDE SELECT B1

SIDE SELECT EVEN PARITY

ENGINE BU DISPLAY ENABLE NO. 1ENGINE BU DISPLAY ENABLE NO. 2 34

190J1BIC-600 IAC NO. 1

PRIMARY RSB

MKR SENSEL

C160J2AV-850A AUDIO CONTROLUNIT NO. 2

AV-850A AUDIO CONTROLUNIT NO. 1

34

C190J1BIC-600 IAC NO. 2

137J1B

(H)

(L)

40

41

DA-800 DAU UNITNO. 2

CC

DD

FF

GG

hi

V

W

M

N

S

P

R

U

Y

c

x

AA

t

C144J1C144J1

(H)

(L)

RM-855 RADIO MANAGEMENTUNIT NO. 2

MKR SENSE LEFTMKR SENSE RIGHT

ATC IDENT

W.O.W.

W.O.W. POLARITY

SIGNAL GROUND

ON/OFF PAGE DISABLE

SIDE SELECT B1

CSDB FMSBUS XMTR

FMSCSDB

(H)

(L)

ARINC 429ENGINE NO. 1

BU DATA

(H)

(L)

ARINC 429ENGINE NO. 2

BU DATA

+28 V DCESSENTIAL POWER

+28 V DCPRIMARY POWER

ESSENTIAL POWERRETURN

PRIMARY POWERRETURN

(H)

(L)

SECONDARYRSB

(H)

(L)

PRIMARYRSB

ENGINE BACKUPDISPLAY ENABLE

ENGINE BACKUPDISPLAY ENABLE

TO AIRCRAFTSIGNAL GROUND


LEFT SIDESECONDARYRSB

RIGHT SIDESECONDARY

RSB

ARINC 429BACKUPENGINE DATA

ARINC 429BACKUP

ENGINE DATA

FMSCSDB

(H)

(L)

ARINC 429ENGINE NO. 1BU DATA

(H)

(L)

ARINC 429ENGINE NO. 2BU DATA

(H)

(L)PRIMARYRSB

(H)

(L)SECONDARYRSB

(H)

(L)

31

30

L CHASSIS GROUND

FROM YOKE MOUNTEDIDENT SWITCHES

AD-49865-R1@

164J1RNZ-851 NAV UNITNO. 1

VOR/ILSCSDB NO. 2

(H)

(L)

B102

A84

E

F

(H)

(L)VOR/ILS BACKUPSERIAL DATA

(Optional)

28 V DC AVIONIC ESSENTIAL BUS

28 V DC BUS 2

DC GROUND

DIMMING BUS

LIGHTING GROUND

C

LL

D

MM

G

H

+5 V LIGHTING

LIGHTING RETURN

+28 V DCESSENTIAL POWER

+28 V DCPRIMARY POWER

ESSENTIAL POWERRETURN

PRIMARY POWERRETURN

LCHASSIS GROUND

C164J1 RNZ-851 NAV UNITNO. 2

VOR/ILSCSDB NO. 2

(H)

(L)

B102

A84

E

F

(H)

(L)VOR/ILS BACKUP

SERIAL DATA

Figure 2-6-20. Radio Management Unit Interface Diagram



B. CD-850 Clearance Delivery Head Interface

See Figure 2-6-21. The NAV ACH SHIFT LOAD and NAV ACH CLOCK inputs come fromthe NAV module and are used to control when the NAV ACH DATA is sent to the NAVmodule. The COM ACH DATA is sent to the COM module on a recurring cycle. Thedifference in timing is because of differences in module software operations.

The WOW input allows the POST to take place only if the aircraft is on the ground whenthe CDH is turned on.

The RSB input is used to display the active frequencies of the COM2 and NAV2 modules.Should there be an in-flight power failure, the CDH will come back on line displaying thesame frequencies that were active prior to the failure.

Other inputs include power, lighting and ground connections.

90

88

69

72

70

58

F

G

f

P

M

N

A BC164J1

165J1

H

L

H

L

H

L

NAV AUXSHIFT/LOAD

NAV AUXCLOCK

NAV AUXDATA

RNZ-851 NAVUNIT NO. 2

33

25

C143J1

H

L

COM AUXDATA

RCZ-851 COMUNIT NO. 2

H

L

H

L

H

L

NAV AUXSHIFT/LOAD

NAV AUXCLOCK

NAV AUXDATA

CD-850 CLEARANCEDELIVERY HEAD (CDH)

H

L

H

L

RSB(LISTEN ONLY)

COM AUXOUTPUT DATA

W.O.W.

H

J

g

R

Y

PRIMARY RSB

C

L

q

W.O.W. SENSE

COM B.W. SELECT

CONFIGURATION COMMON

E

j

D

Z

H

L

H

L

+28 V INPUTPOWER

+5 V DC LIGHTING

POWER GROUND

PILOT'S DIMMING BUS

LIGHTING GROUND

165J1 28 V DCDC BUS 2


AD-49871-R1@

Figure 2-6-21. CD-850 Clearance Delivery Control Head Interface Diagram



C. AV-850A Audio Control Unit Interface

See Figure 2-6-22.

The AV-850A Audio Control Unit receives digitized audio from the remote units via twohigh-speed digital audio buses. Each audio control unit then selects the appropriatechannels from this digital audio bus, and reconstitutes headphone and speaker signals.This allows the system to provide the capability to individually select the radio function thateach crew member desires to hear. There is a row of microphone selector buttons alongthe top edge, which when pressed, will select the desired transmitter/receiver and at thesame time automatically enable the receive audio associated with that transmitter/receiver,regardless of the condition of the audio on/off buttons. The audio on/off buttons arelocated on the lower rows of the audio control unit. Pushing the button will cause it to latchstowed and the audio associated with that button will be turned off. Pushing the buttonagain will allow it to pop out and energize the audio into the speaker and headphone andallow the audio level to be adjusted by rotating the button. Also included are mastervolume controls for both speaker and headphone.

The audio control unit also has numerous connections for intercom, crew annunciation,crew communication, hot microphone, passenger address, etc., and full-time emergencywarning inputs from aircraft systems. Cross-cockpit audio is provided so that pilot andcopilot will remain coordinated with each other in their selection and use of the radiosystem components.

The Embraer 145 also has an AV-850A installed at the observer’s station.

Digital audio offers the advantage of complete independence from grounding problemswithin the aircraft and the absolute elimination of ground noise pick-up, whine, andcross-talk. Having the audio digitized also offers the advantage that when recoveringthe analog information from the digital, each volume control may be independently setby each crew member. For instance, the pilot may wish to have COM1 very loud andCOM2 very soft in his headset while the copilot desires the VOR to be loud, COM1 to besoft and COM2 to be moderate. This is easily set at the audio control units simply byadjusting the volumes to their own desire. Also, by having the audio system digitized,various filtering and priority functions can be easily accomplished to enhance the operationand the human interfaces.

The audio control unit has provisions for a combination of emergency operations. In thecase of power down or failure of the audio system, there is one location (in the upperright-hand corner of the audio control unit) where pushing the microphone select (labeledEMER) for the emergency COM bypasses all the circuitry within the audio control unit andplaces the emergency COM and NAV audio into the headphone circuitry. Emergency audiois analog audio from the COM and NAV modules that are connected to the CD-850 CDH.This function is also convenient during ground operation when minimum power usage isdesired.



See Figure 2-6-22 (sheet 1) for the following:

• Microphone HI, LO and PTT inputs and outputs (except the maintenance loop)

• Interphone audio and PTT

• Emergency analog audio inputs from COM2 and NAV2

• Speaker outputs

• Cockpit Voice Recorder (CVR) outputs

• Flight Data Recorder (FDR) PTT discretes via the pilot’s DAU

• Other inputs include power, lighting and ground connections.

See Figure 2-6-22 (sheet 2) for the following:

• Passenger Address System interconnections

• Interphone XMIT

• Warning audio inputs

• Digital audio bus

• The receive audio from the HF transceiver is applied to the copilot’s COM unit clustermodule for access to the digital audio system.

• The receive audio from VHF COM No. 3 is applied to the pilot’s COM unit clustermodule for access to the digital audio system.

See Figure 2-6-22 (sheet 3) for all of the above as it applies to the observer’s positionaudio control unit.



(Blank Page)



Y

L

t

k

K

J

A

h

160J1

160J2

TO LEFT SIDE COCKPIT SPEAKER

TO PILOT'S HEADPHONE JACK

LEFT SIDE HAND MIC JACK

LEFT YOKE MOUNTED PTT BUTTON

LEFT SIDE OXYGEN MASK

LEFT SIDE BOOM MIC JACK

TO COCKPIT VOICE RECORDER

LEFT YOKE MOUNTED

INTERPHONE BUTTON

E

Y

V

r

p

(H)

(L)

HAND MIC PTT

MASK/BOOM PTT

MASK MIC HI

BOOM MIC HI

(H)

(L)

(H)

(L)

(H)

(L)

COCKPIT SPEAKER

PHONE

HAND MIC AUDIO

VOICE RECORDER

U

V

M

q

g

160J1

+28 V DC ESSENTIAL BUS

DC GROUND

PILOT'S DIMMING BUS

LIGHTING GROUND

+28 V DC (1)

+28 V DC (2)

+28 V DC RETURN

LIGHTING RETURN

+5 V DC LIGHTING

T

S

X

a

b

Z

(H)

(L)

(H)

(L)

HF MIC AUDIO

HF PTT

COM 3MIC AUDIO

COM 3 PTT

s

j

D

W

C

P

H

B

R

E

F

G

f

m

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

VHF NO. 1 MIC AUDIO

EMERGENCY COM PTT

INTERPHONE AUDIO

VHF NO. 2 MIC AUDIO

VHF NO. 2 PTT

EMERGENCY COM MIC AUDIO

VHF NO. 1 PTT

EMERGENCY COM AUDIO HI

EMERGENCY NAV AUDIO HI

EMERGENCY NAV/COM AUDIO LO

160J1

92

106

21

37

143J1

(H)

(L)

PTT

HF COM TRANSMIT

MIC AUDIOINPUT

RCZ-851E COM UNIT NO. 1

AV-850A AUDIO CONTROL UNIT NO. 1

T

S

X

a

b

Z

s

j

D

W

C

P

H

B

R

E

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

C160J1

HF MIC AUDIO

HF PTT

VHF NO. 2 PTT

EMERGENCY COM PTT

INTERPHONE AUDIO

VHF NO. 2 MIC AUDIO

EMERGENCY COM MIC AUDIO

VHF NO. 1 MIC AUDIO

F

G

f

m

VHF NO. 1 PTT




(H)

(L)

92

106

21

37

C143J1

HF COM TRANSMIT

PTT

MIC AUDIOINPUT


AV-850A AUDIO CONTROL UNIT NO. 2

TO RIGHT SIDE COCKPIT SPEAKER

TO COPILOT'S HEADPHONE JACK

RIGHT SIDE HAND MIC JACK

RIGHT YOKE MOUNTED PTT BUTTON

RIGHT SIDE OXYGEN MASK

RIGHT SIDE BOOM MIC JACK

RIGHT SIDE MASK/BOOM LO


RIGHT YOKE MOUNTEDINTERPHONE BUTTON

+28 V DC BUS

DC GROUND

COPILOT'S DIMMING BUS

LIGHTING GROUND

Y

L

t

k

K

J

A

h

C160J1

C160J2

E

Y

V

rp

(H)

(L)

HAND MIC PTT

MASK/BOOM PTT

MASK MIC HI

BOOM MIC HI

(H)

(L)

(H)

(L)

(H)

(L)

COCKPIT SPEAKER

PHONE

HAND MIC AUDIO

VOICE RECORDER

U

V

M

q

g

C160J1

+28 V DC (1)

+28 V DC (2)

+28 V DC RETURN

LIGHTING RETURN

+5 V DC LIGHTING

TO COM 3

TO HF TRANSCEIVER

INTERPHONE PTT INTERPHONE PTT

LEFT SIDE MASK/BOOM LO

37

136J2ADA-800 DAU NO. 1

41

136J1A

68

136J1B

69

(H)

(L)FLIGHT DATAREC OUT

TO LEFT SIDE CABIN SPEAKERe CABIN SPEAKER

D COMMON GROUNDLEFT YOKE MOUNTED PTT BUTTON

COM 3MIC AUDIO

COM 3 PTT

DCOMMON GROUNDLEFT YOKE MOUNTED PTT BUTTON

TO RIGHT SIDE CABIN SPEAKEReCABIN SPEAKER

AD-49873@

31

29

(H)

(L)ANALOGAUDIO

B36

B51

164J1

(H)

(L)

VOR/LOCANALOG AUDIO

RNZ-851 NAV UNIT NO. 2

COPILOT RADIOXMTR KEYING

PILOT RADIOXMTR KEYING

TO FLIGHT DATA RECORDER

Figure 2-6-22 (Sheet 1). Audio Control Unit Interface Diagram



(H)

(L)

COM NO. 3

AUDIO OUTPUT

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

PA MIC

PA MIC PTT

STEREO MUTE

i

U

j

M

f

P

m

J

q

e

k

MAINT PHONE

MAINT 1 MIC

MAINT 1 ENABLE

WARNING3 AUDIO

160J2

PART OFPASSENGER

ADDRESSSYSTEM

AMPLIFIER

i

U

j

M

(H)

(L)

(L)

(H)

PA MIC PTT

STEREO MUTE

PA MIC

WARNING3 AUDIO

c cINTERPHONEXMIT

INTERPHONEXMIT

c

n

d

p

c

n

d

p

160J1 C160J1

(L)

(H)

(L)

(H)

(H)

(L)

(H)

(L)

DIGITALAUDIOBUS NO. 2




(L)

(H)

(L)

(H)

(H)

(L)

(H)

(L)

56

70

B42

B29

56

70

B42

B29DIGITAL AUDIO BUS

DIGITALAUDIO BUS

164J1

143J1

C164J1

C143J1

DIGITAL AUDIO BUS

DIGITALAUDIO BUS





AV-850A AUDIO CONTROLUNIT NO. 2

AV-850A AUDIO CONTROLUNIT NO. 1 C160J2

PTT

143J1

66

50

37

(H)

(L)

COM 3AUDIO IN

COM 3 PTT

C143J1

66

50

37

(H)

(L)

HF NO. 1

AUDIO OUTPUT

PTT

(H)

(L)

HF AUDIO IN

HF PTT

m

J

(H)

(L)

AURAL WARNINGUNIT

AUDIO OUTPUT

160J2

W

B

(H)

(L)

WARNING1 AUDIO

(H)

(L)

AURAL WARNINGUNIT

AUDIO OUTPUT

C160J2

W

B

(H)

(L)

WARNING1 AUDIO

AD49866@




t

k

h

E160J1

TO OBSERVER'S HEADPHONE JACK

OBSERVER'S PTT BUTTONMASK/BOOM PTT

(H)

(L)PHONE T

S

X

(H)

(L)HF MIC AUDIO

HF PTT

s

j

D

W

C

R

E

F

G

f

m

(H)

(L)

(H)

(L)

(H)

(L)

VHF NO. 1 MIC AUDIO

INTERPHONE AUDIO

VHF NO. 2 MIC AUDIO

VHF NO. 2 PTT

VHF NO. 1 PTT




E160J1AV-850A AUDIO CONTROL UNIT (OBSERVER'S)

CONNECTED TO THESAME CONNECTORSAND PINS ON BOTHPILOT'S AND COPILOT'SAUDIO CONTROL UNITS

K

J

A

q

g

+28 V DC NON-ESSENTIAL BUS 1

DC GROUND

LIGHTING GROUND

+28 V DC (1)

+28 V DC (2)

+28 V DC RETURN

LIGHTING RETURN

+5 V DC LIGHTINGLIGHTING GROUND

E160J2

OBSERVER'S OXYGEN MASK

OBSERVER'S BOOM MIC JACK

OBSERVER'S MASK/BOOM LO


OBSERVER'S MOUNTED

INTERPHONE BUTTON

E

Y

V

r

p

MASK MIC HI

BOOM MIC HI

(H)

(L)VOICE RECORDER

INTERPHONE PTT

C SPARE GROUND

c

n

d

p

(H)

(L)

(H)

(L)



E160J2

cINTERPHONE

XMIT

COMMON GROUNDDOBSERVER'S PTT

(H)

(L)HAND MIC AUDIO

HAND MIC PTT

OBSERVER'S HAND MIC JACKa

Z

b

(H)

(L)VHF NO. 3 MIC AUDIO

VHF NO. 3 PTT

VHF COM NO. 3(H)

(L)MIC AUDIO

PTT

a

Z

(H)

(L)

WARNING1 AUDIO

AURAL WARNINGUNIT

(H)

(L)AUDIOOUTPUT

U

V

M

Y

L

TO OBSERVER'S CABIN SPEAKER

TO OBSERVER'S COCKPIT SPEAKER

e CABIN SPEAKER(H)

(L)COCKPIT SPEAKER

(H)

(L)

(H)

(L)

(H)

(L)

PA MIC

PA MIC PTT

STEREO MUTE

i

U

j

M

m

J

g

n

R

MAINT 2 MIC

MAINT 2 ENABLE

WARNING3 AUDIO

PART OFPASSENGER

ADDRESSSYSTEM

AMPLIFIER

AD-49872@




D. Communications Unit Interface

See Figure 2-6-23.

Connections shown include power, ground, and RSB, all of which have been covered inprevious sections.

Command data for the TCAS computer unit originate at the RMU and are transmitted viaRSB. The cluster modules in each COM UNIT receive the data and pass it to the RMUselected active transponder, which in turn sends the data to the TCAS CU via ARINC 429.TCAS reply coordination data is handled in the reverse.

During a cold start power up, the cluster board sends word load and clock to the strapoption board and receives data back. This data is then sent to the modules via RCB.

The rack mounted fan is controlled by the cluster module and is turned on when anymodule reports a temperature above the threshold. The threshold is different in eachmodule. When the temperature drops below the threshold, the cluster module turns the fanoff. The fan is not required for system operation.

Encoded altitude data is transmitted by each MADC via ARINC 429 to both transponders.Each transponder is programmed to recognize the on-side MADC as the normal source ofdata. Should there be failure in an MADC, pressing the MADC REVERSION switch on thesame side will latch the switch, and, in turn, will send a ground to the same sidetransponder causing it to accept encoded altitude data from the cross-side MADC.

The SIMULCOM line goes to ground when either VHF COM transceiver is transmitting.This connection allows communication on different frequencies simultaneously withminimum interference.

The WOW is used by the transponder to disable Mode S replies when the aircraft is on theground.

The mutual suppression bus goes to ground when any of the transmitters on the line istransmitting. This prevents the DME, TCAS, and transponders from interfering with eachother during otherwise normal operation.

Whichever transponder has been selected as the active transponder sends a grounddiscrete to the aircraft antenna relay. Therefore, only one set of transponder antennae arerequired to be mounted on the aircraft.



(Blank Page)



4

5

6

8

9

17

18

19

7

20

RACK MOUNTED STRAPOPTION BOARD NO. 1(P/O MT-852)

WORD LOAD

CLOCK

DATA

+5 V DC POWER

POWER RETURN

2

6

7

3

4

(ORANGE)

(BLUE)

(WHITE)

(RED)

(BLACK)


35

67

10

11

74

61

83

55

87

90

104

(H)

(L)

(H)

(L)

ARINC 429TX CO-ORD OUTPUT

ARINC 429TX CO-ORD INPUT

(H)

(L)PRIMARY RSB

(H)

(L)SECONDARY RSB


14F

14G

15J

15K

(H)

(L)

(H)

(L)

ARINC 429XT CO-ORD INPUT

ARINC 429XT CO-ORD OUTPUT

2

54

(RED)

(BLUE)

RACK MOUNTED FANNO. 1 (P/O MT-852)

+28 V DC FAN POWER

POWER RETURN

SIMULCOMPOWER RETURN ATC

CHASSIS GROUND

SIMULCOM

CHASSIS GROUND

102

88

101

16

43

STRAP ASSY WORD LOAD

STRAP ASSY CLOCK

STRAP ASSY SERIAL DATA

STRAP ASSY POWER

STRAP ASSY POWER RETURN

94

83

94

POWER RETURN ATC

POWER RETURN COM

POWER RETURN COM

POWER RETURN COM

+28 V DC ATC

+28 V DC ATC

+28 V DC COM

+28 V DC COM

+28 V DC COM

4

5

6

8

9

17

18

19

7

20POWER RETURN ATC

POWER RETURN ATC

POWER RETURN COM

POWER RETURN COM

POWER RETURN COM

+28 V DC ATC

+28 V DC ATC

+28 V DC COM

+28 V DC COM

+28 V DC COM+28 V DC AMDC BUS 2A

POWER GROUND

+28 V DC ESSDC BUS 1

POWER GROUND

2

6

7

3

4

2

54

102

88

101

16

43

RACK FAN POWER

GROUND TO ENABLE FAN


STRAP ASSY CLOCK


STRAP ASSY POWER


RACK FAN POWER



RIGHT SIDESECONDARY

RSB


35

67

10

11

74

61

FROM AIRCRAFT WEIGHT ON WHEELS

(ORANGE)

(BLUE)

(WHITE)

(RED)

(BLACK)

(RED)

(BLUE)

WORD LOAD

CLOCK

DATA

+5 V DC POWER

POWER RETURN

+28 V DC FAN POWER

POWER RETURN



ATC MUT SUP

WOW

ATC MUT SUP

WOW

PRIMARY RSB

SIMULCOM

143J1 C143J1

(H)

(L)PRIMARY RSB

(H)

(L)SECONDARY RSB


C143J1

143J1

193RMP

TO SUPPRESSION BUS

R.F. ANTENNA CABLES AND R.F. CONNECTORS NOT SHOWN.NOTE:

24 24SIMULCOM RTN SIMULCOM RTN

ARINC 429NO. 3 OUT

AZ-850 MADC NO. 2(H)

(L)

(H)

(L)66

67

66

67

ARINC 429NO. 3 OUT

AZ-850 MADC NO. 1(H)

(L)

(H)

(L)66

67

66

67

99

86

99

86

100

69

10069

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)PRI MADC

429 IN

PRI MADC429 IN

SEC MADC429 IN

SEC MADC429 IN

9J1

C9J1

9J1

C9J1

MADC REV

DC-550 NO. 2

74

C115J1

MADC REV

DC-550 NO. 1

74

115J1

ALT SRC NO. 2 SELALT SRC NO. 2 SEL 8080

1 2

MADC REVERSION (LATCHING)

55

87

90

104

(H)

(L)

(H)

(L)


ARINC 429TX CO-ORD INPUT


14H

14J

14A

14B

(H)

(L)

(H)

(L)


ARINC 429XT CO-ORD OUTPUT

193RMP

XPDR ACTIVE XPDR ACTIVE9191 AIRCRAFTANTENNA

RELAY

28V DC AMDC BUS 1B

28V DC AMDC BUS 2B

AD-49867-R1@

MUTUAL SUPPRESSION IN 12

193LBPFROMMUTUALSUP. BUS

Figure 2-6-23. Communication Unit Interface Diagram



E. Navigation Unit Interface

See Figure 2-6-24.

Connections shown include power, ground, and RSB, all of which have already beendiscussed.

Connections between RCB to MLS and MLS to RCB instruct the cluster module that thereis no MLS receiver on board, and therefore, the cluster module will not file failure tocommunicate messages.

The ADF antenna receives power and modulation signals from the ADF receiver via thecluster module. The signal from the antenna is via triaxial cable, not shown on thisdiagram. The ADF antenna contains the self-test signal generator which is used duringself-test procedures.

During a cold start power up, the cluster board sends word load and clock to the strapoption board and receives data back. This data is then sent to the modules via RCB.

The rack-mounted fan is controlled by the cluster module and is turned on when anymodule reports a temperature above the threshold. The threshold is different in eachmodule. When the temperature drops below the threshold, the cluster module turns the fanoff. The fan is not required for system operation.

The HF PTT key line is used by the ADF receiver to prevent the ADF from pointing to theHF antenna during transmissions. The ADF pointer will "freeze" at the last position forabout 10 seconds. After that, the pointer will move to 90 degrees relative to the nose ofthe aircraft and remain there until HF transmission ceases. At that time, the pointer willreturn to the "freeze" position and will then move to the appropriate new position withoutany sudden movements.

The mutual suppression bus goes to ground when any of the transmitters on the line istransmitting. This prevents the DME, TCAS, and transponders from talking to each otherduring otherwise normal operation.

The optional FMS computer receives both VOR radial information and DME distanceinformation from both NAV units. The FMS computer determines which data to use. Formore complete details on the FMS, see Chapter 8, this section, this manual.



(Blank Page)



164J1RNZ-851 NAV UNIT NO. 1

AT-860 ADF ANTENNA(NO. 1)

H

B

K

C

G

J

A104

A91

A101

A77

B1

B27

ADF SIN MOD

ADF COS MOD

ADF LOOP ENABLE

ADF ANTENNA TEST

ADF +15 V DC POWER

ADF POWER RETURN

ADF SIN MOD

ADF COS MOD

ADF LOOP ENABLE

ADF ANTENNA TEST

ADF ANTENNA POWER

ADF ANTENNA POWER RETURN

(ORANGE)

(BLUE)

(RED)

(BLACK)

(WHITE)

2

6

7

3

4

B84

B59

B58

B40

B26

B2

B44

(RED)

(BLUE)

WORD LOAD

CLOCK

DATA

+5 V DC POWER

POWER RETURN


STRAP ASSY CLOCK


STRAP ASSY POWER


RACK FAN POWER


+28 V DC FAN POWER

POWER RETURN



164J1

A85

A98

B39

B52

A86

A99

A87

A100

B38

B50

A83

A95

A94

B25

A103

B98

+28 V DC VOR/ILS

+28 V DC VOR/ILS

+28 V DC ADF

+28 V DC DME

+28 V DC VOR/ILS

+28 V DC VOR/ILS

+28 V DC DME

+28 V DC ADF

POWER RETURN VOR/ILS

POWER RETURN DME

POWER RETURN ADF

POWER RETURN CLUSTER


POWER RETURN DME

POWER RETURN ADF


DC GROUND

28 V DCAM DC BUS 1A

28 V DC AM ESSDC BUS 1

A66ADF HF KEY

FROM HF PTT

A42

TO SUPPRESSION BUS

DME MUT SUP

B71

B81

(H)

(L)PRIMARY RSB

PRIMARY RSB

A88

A102

(H)

(L)SECONDARY RSB

A57

A44

B89

A59

121J1B

(H)

(L)

(H)

(L)

DME CSDBOUTPUT

VOR/ILS CSDBOUTPUT

(L)

(H)

CSDB RCVRNAV PRI

NZ-2000 NAVIGATION COMPUTER(OPTIONAL)

NOTE: R.F. ANTENNA CABLES AND

R.F. CONNECTORS NOT SHOWN.

A85

A98

B39

B52

A86

A99

A87

A100

B38

B50

A83

A95

A94

B25

A103

B98

A66

A42

B71

B81

A88

A102

+28 V DC VOR/ILS

+28 V DC ADF

+28 V DC DME

+28 V DC VOR/ILS

+28 V DC VOR/ILS

+28 V DC DME

+28 V DC ADF


POWER RETURN DME

POWER RETURN ADF



POWER RETURN DME

POWER RETURN ADF


ADF HF KEY

DME MUT SUP

(H)

(L)

(H)

(L)

PRIMARY RSB

SECONDARY RSB

RNZ-851 NAV UNIT NO. 2C164J1 C164J1

A104

A91

A101

A77

B1

B27

B84

B59

B58

B40

B26

B2

B44GROUND TO ENABLE FAN

RACK FAN POWER


STRAP ASSY CLOCK


STRAP ASSY POWER


ADF SIN MOD

ADF COS MOD

ADF LOOP ENABLE

ADF ANTENNA TEST

ADF ANTENNA POWER

AT-860 ADF ANTENNA(NO. 2)

H

B

K

C

G

J

ADF SIN MOD

ADF COS MOD

ADF LOOP ENABLE

ADF ANTENNA TEST

ADF +15 V DC POWER

ADF POWER RETURN

(ORANGE)

(BLUE)

(RED)

(BLACK)

(WHITE)

2

6

7

3

4

(RED)

(BLUE)

WORD LOAD

CLOCK

DATA

+5 V DC POWER

POWER RETURN

+28 V DC FAN POWER

POWER RETURN


C158J1158J1

DC GROUND

RACK MOUNTED STRAP OPTION BOARD NO. 2(P/O MT-851)

A21

A34

A25

A38

(H)

(L)

(H)

(L)

RCB TO MLS

(IF NO MLS)

RCB FROM MLS

ADF ANTENNA POWERRETURN


RIGHT SIDESECONDARY

RSB

8296

(L)

(H)

CSDB RCVRDME PRI

A57

A44

B89

A59

(L)

(H)

(L)

(H)CSDB RCVR

NAV SEC

CSDB RCVRDME SEC

DME CSDBOUTPUT

VOR/ILS CSDBOUTPUT

81

80

74

73

95

94

(H)

(L)

(H)

(L)

28 V DC AMDC BUS 2A

28 V DC AMDC BUS 2B

+28 V DC VOR/ILS

AD-49868-R1@

(H)

(L)

(H)

(L)

A21

A34

A25

A38

RCB TO MLS

RCB FROM MLS

(IF NO MLS)

Figure 2-6-24. Navigation Unit Interface Diagram



4. Fault Monitoring

Fault indications are presented on the RMU and on the PFD display tube. See Figures 2-6-25and 2-6-26.

A. PFD Indications

Loss of valid vertical deviation from the NAV receiver will cause the following to happen:

• Removal of the vertical deviation pointer

• Scale to be red X’d.

Loss of valid lateral deviation from the NAV receiver will cause the following to happen:

• Removal of the HSI lateral deviation pointer

• HSI lateral deviation scale to be red X’d.

Loss of valid distance information from the DME module will cause the following to happen:

• Removal of the morse identifier

• Green dash of the distance digital readout

• Amber dash of the groundspeed digital readout

• Amber dash of the time to go digital readout.

Loss of valid bearing information from the NAV receiver will cause the following to happen:

• Removal of the HSI lateral deviation pointer

• HSI lateral deviation scale to be red X’d

• Removal of the TO/FROM display

• Removal of the absolute bearing pointers.

B. RMU Indications

Any failure of a module will cause the RMU to remove the frequencies or operatingcommands associated with that particular function and replace them with dashes.

Should a transponder that is operating in the standby mode fail while the other transponderis active, a red ATC1 INOP or ATC2 INOP message will appear on the bottom line in thetransponder window on the RMU.

The radio system POST and PAST are further described in SECTION 7 of this manual.



19

200

18185 E

170

2160

150

220

HDG ASELAP

10 10

20 20

10 10

20 20

104 0010500

10020

8000

29.92 IN

CRS360

VOR1

ADF2

HDG360

10000

1

23

32

1

N

6

15

24

S

W

E

GSPD- - - KTS

- - -. - NM

IAS

AD-51616@

Figure 2-6-25. PFD Radio System Failure Indications



AD-48976@

Figure 2-6-26. RMU Failure Indications



(Blank Page)





2.7 TRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM (TCAS II) . . . . . . . . . . . . 2-7-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-1


A. RT-910 TCAS Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-2

B. AT-910 Directional Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-4

C. Typical Bottom Omnidirectional Antenna . . . . . . . . . . . . . . . . . . . . 2-7-5

D. Other Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-5

E. TCAS/MFD Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-6

F. TCAS/RMU Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-6

(1) Line Select Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-6

(2) CODE Line Select Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-6

(3) MODE Line Select Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-7

(4) RANGE Line Select Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-8

(5) SURVEILLANCE WINDOW Line Select Key . . . . . . . . . . . . . . 2-7-8

(6) PGE Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-8

G. TCAS Mode Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-10

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-11

A. TCAS Computer Unit Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-12

B. TCAS ARINC 429 Output Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-13

C. TCAS Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-17

(1) Traffic Advisories . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-17

(2) Resolution Advisories . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-20

D. TCAS Display Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-22

(1) Traffic Advisory (TA) Logic . . . . . . . . . . . . . . . . . . . . . . . . 2-7-22

(2) Resolution Advisory (RA) Logic . . . . . . . . . . . . . . . . . . . . . 2-7-22

E. TCAS Preflight Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-22

(1) Activate TCAS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-22

(2) Activate TCAS Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-22

(3) Perform Self-test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-23



B. Fault Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-24




FIGURE/TITLE PAGE


Figure 2-7-1. RT-910 TCAS Computer Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-2

Figure 2-7-2. AT-910 Directional Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-4

Figure 2-7-3. Typical Bottom Omnidirectional Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-5

Figure 2-7-4. Typical RM-855 Radio Management Unit (RMU) . . . . . . . . . . . . . . . . . . . . . . 2-7-7

Figure 2-7-5. TCAS Computer Unit Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . 2-7-15/16



Figure 2-7-8. TCAS PFD Symbology . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-21

List of Tables

TABLE/TITLE PAGE

Table 2-7-1. RT-910 TCAS Computer Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . 2-7-2

Table 2-7-2. AT-910 Directional Antenna Leading Particulars . . . . . . . . . . . . . . . . . . . . . 2-7-4

Table 2-7-3. RT-910 TCAS Computer ARINC 429 Output Data Table . . . . . . . . . . . . . . . . 2-7-13

Table 2-7-4. RT-910 TCAS Computer-To-Mode S Transponder Data Table . . . . . . . . . . . 2-7-14

Table 2-7-5. RCZ-851E Communications Unit-To-TCAS Computer Data Table . . . . . . . . 2-7-14

Table 2-7-6. MFD/TCAS Symbology . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-7-17



SECTION 2.7TRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM (TCAS II)

1. General

The PRIMUS® 1000 Integrated Avionics System installed in the Embraer 145 aircraft includes aHoneywell traffic alert and collision avoidance system (TCAS), a TCAS II system, which iscapable of providing conflict resolution advisories in the form of vertical guidance. The TCAS IIsystem can be operated in several modes to display traffic advisories (TAs) and/or resolutionadvisories (RAs).

The TCAS II system consists of:

• One RT-910 TCAS Computer Unit

• One top-mounted AT-910 TCAS Directional Antenna

• One bottom-mounted Omnidirectional Antenna (or)

• One optional bottom-mounted AT-910 TCAS Directional Antenna.

The TCAS II system also interfaces with:

• One or two Diversity Mode S Transponder(s) with Antennas


• Two Radio Management Units (RMUs)

• Aircraft Audio System.

The transponders, the RMUs, and the audio system are part of the PRIMUS® II Integrated RadioSystem and full operating details of the RMU are described in SECTION 2, CHAPTER 6.

For full operational information on the TCAS, refer to Honeywell Pub. No. 28-1146-70, TrafficAlert and Collision Avoidance System (TCAS) Pilot’s Manual.




A. RT-910 TCAS Computer

Figure 2-7-1 shows a graphical view of the RT-910 TCAS Computer. It is located in theavionics nose bay of the aircraft. Table 2-7-1 provides items and specifications that areparticular to the computer.

Table 2-7-1. RT-910 TCAS Computer Leading Particulars

AD-32943@

Figure 2-7-1. RT-910 TCAS Computer Unit

Item Specification

Dimensions (maximum) . . . . . . . . . . . . . . . . 6 MCU short per ARINC 600

• Height . . . . . . . . . . . . . . . . . . . . . . . . . 7.64 in. (194.06 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . 7.75 in. (196.85 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . 12.75 in. (323.85 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24.0 lb (10.9 kg)

Power Requirements . . . . . . . . . . . . . . . . . . 115 V ac, 85 W (max)

User Replaceable Parts . . . . . . . . . . . . . . . . None

Mating Connector . . . . . . . . . . . . . . . . . . . . TRI-Star Part No. C-06B5-9901-0100

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . Barry Series 95578



The TCAS computer unit (CU) contains the RF transmitter and receivers necessary tointerrogate and receive replies from transponder equipped aircraft. Dual microprocessorsare utilized to implement the surveillance and collision avoidance algorithms to decidewhether an intruder aircraft should be considered a threat and then to determine theappropriate vertical response to avoid a midair collision or near midair incident. Inaddition, output data is provided to drive displays and the aircraft audio system to informthe flight crew as to what action to take or avoid. An interface is provided with an on-boardMode S transponder in order to coordinate avoidance maneuvers with otherTCAS-equipped aircraft.



B. AT-910 Directional Antenna

Figure 2-7-2 shows a graphical view of the AT-910 Directional Antenna and Table 2-7-2provides items and specifications that are particular to the antenna.

AD-32826@

Figure 2-7-2. AT-910 Directional Antenna

Table 2-7-2. AT-910 Directional Antenna Leading Particulars

Item Specification


• Height (Outside Aircraft) . . . . . . . . . . . . 0.806 in. (20.47 mm)

• Height (Inside Aircraft) . . . . . . . . . . . . . 1.56 in. (39.62 mm)

• Diameter . . . . . . . . . . . . . . . . . . . . . . . 9.31 in. (236.47 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.80 lb (1.30 kg)

User Replaceable Parts . . . . . . . . . . . . . . . . None

Mating Connectors . . . . . . . . . . . . . . . . . . . Type TNC

The AT-910 Directional Antenna (located on top of the aircraft fuselage) in conjunction withfour receivers in the TCAS computer unit provide the capability to determine the bearing ofthe intruder. Since TCAS II is a vertical-only system, intruder bearing is not used in thecomputation of the escape or limit maneuver. Intruder bearing is used only to enable theflight crew to more easily locate the intruder visually. The TCAS II will also accommodatea bottom-mounted directional antenna if so desired by the user. Such an installation willslightly increase the areas in which bearing information is available. The AT-910Directional Antenna is connected to the TCAS computer unit by four coaxial cables.

The antenna is capable of receiving replies from all directions simultaneously with bearinginformation using amplitude-ratio monopulse techniques. Insertion loss differences incoaxial cable lengths from the antenna to the TCAS computer need only be matched towithin 0.5 dB which corresponds to a 5- to 10-foot difference in length depending on thespecific cable type. Losses between the antenna and the computer unit must be 2.5 ± 0.5dB, including line connections.



C. Typical Bottom Omnidirectional Antenna

Figure 2-7-3 shows a graphical view of a Typical Bottom Omnidirectional Antenna.

AD32827@

Figure 2-7-3. Typical Bottom Omnidirectional Antenna

The leading particulars for the bottom omnidirectional antenna must be obtained from theselected antenna manufacturer. The omnidirectional antenna is located on the bottom ofthe aircraft fuselage and does not provide any directional information, but does enable theTCAS computer unit to interrogate and receive replies from intruder aircraft locatedbeneath own aircraft. Since TCAS II is a vertical-only system, intruder bearing is not usedin the computation of the escape or limit maneuver.

The omnidirectional antenna must exhibit 50-ohm impedance, and must also exhibit50-ohm fixed resistance to ground. This is necessary for self-test. This omnidirectionalantenna may be identical to that used for the top and bottom antennas connected to theMode S transponder.

D. Other Components

In order for the TCAS to make the flight crew aware of potential threats, a number ofadditional LRUs are used. Their functions are:

• Aural warnings are sent through the aircraft audio system, and to a dedicated TCASspeaker

• Traffic Advisories (TAs) are sent to the IC-600 IACs to be displayed on the MFDs

• Resolution Advisories (RAs) are sent to the IC-600 IACs to be displayed on the PFDVSI display

• Altitude data originates in the MADCs, is sent to the transponders, and then passes tothe TCAS via ARINC 429

• TCAS mode, range, and vertical surveillance window mode control is through the RMU

• Command data from the RMU passes through the diversity Mode S transponder to theTCAS computer unit.



E. TCAS/MFD Controls

The TCAS display on the MFD is controlled by the MFD menu. When the MAIN menu isdisplayed, pressing the button beneath the TCAS legend will toggle the TCAS zoomwindow on and off. TCAS modes are displayed anytime TCAS is selected for display.Selection of TCAS display ON will override any other displays in the area used by thezoom window.

The TCAS display may further be selected for either full time display, or "pop-up" display.In the pop-up mode, there is no display unless a traffic advisory or resolution advisory hasoccurred. See paragraph 2.F.(6)(a)2 for specific procedures.

Whenever TCAS is selected for display on the MFD, a white range arc is displayed, withthe range, in nautical miles, at the right end of the arc. The display range is selected onthe RMU. Available display ranges are 6, 12, 20, and 40 NM. For 6 or 12 mile selections,a ring of 12 small circles (or dots) is displayed around the own aircraft symbol. The ring isat a radius of 2 nautical miles, and at the clock hour positions When the selected range is20 or 40 NM, the 2-mile range ring is replaced by a half range arc in addition to thenumbered range arc as described above.

F. TCAS/RMU Controls

The following paragraphs describe each RMU control that is used during TCAS operation(see Figure 2-7-4.)

(1) Line Select Key

The first push of any line select key moves the yellow cursor to surround the datafield associated with that particular line select key. This then electronically connectsthat data field to the tuning knobs so that the mode or code may be changed.

(2) CODE Line Select Key

Press this key to place the cursor around the transponder code data line. Now thelarge outer tuning knob controls the left two digits, and the smaller inner knobcontrols the right two digits.

Since only one transponder can operate at a time, both RMUs will be displaying thesame transponder information. Therefore, if a code or mode is changed on oneRMU, the other RMU will track it. Since the other RMU is being tuned by a remotesource, the data changed will appear in yellow.

Press and hold this key for more than 2 seconds to change the code to that whichwas stored in the memory. To store a code in memory, dial the desired code withinthe cursor and then press the STO key. The stored code will remain in memoryduring recall and during power down.



(3) MODE Line Select Key

Pressing this key moves the cursor to the mode line and enables several functions.

• Press this key again to toggle between standby and the last active mode.

• With the cursor at the mode line and an active mode selected, press the 1/2 keyto toggle between transponder 1 active and transponder 2 active.

• Twist either tune knob to change the active mode. The available selections are:ATC ON, ATC ALT, TA ONLY, TA/RA.

• If the system does not include two diversity transponders, the TCAS will only beactive when the diversity transponder is selected as active. In this case, the 1/2button will also cause the banner line to change from ATC/TCAS (when thediversity transponder is selected as active) to ATC (when the non-diversitytransponder is selected as active).

Additionally, when the non-diversity transponder is selected to active and the bannerline shows ATC, only the ATC ON and ATC ALT operational modes may be selected.

The third data line in this window displays either the Flight ID number or errormessages, as appropriate.

TUNESQ

ID

DIM

PGE

1/2

TST

STO

DME

COM 1 NAV1

MEMORY-3

ATC/TCAS ADF1

1 TA/RA

RANGE:

123.20131.27

1471

NORMAL

MEMORY-1

110.30

109.35

ADF

162.5

6

CURSOR

FUNCTIONKEYS

TRANSFER("FLIP-FLOP")KEY

LINE SELECTKEY

TUNINGKNOBS

PHOTO SENSOR

AD-51053@

TCAS DSPY 1

Figure 2-7-4. Typical RM-855 Radio Management Unit (RMU)



(4) RANGE Line Select Key

Press this key to move the cursor to this line. Press this key again or twist eithertuning knob to select the TCAS display range. Available selections are 6, 12, 20,and 40 NM. The 2-mile range ring will be displayed proportional to the rangeselected. When the range is 20 or 40 NM, the 2-mile range ring is removed and ahalf-range arc will be added within the full range arc.

(5) SURVEILLANCE WINDOW Line Select Key

Press this key to move the cursor to this line. Press this key again or twist eithertuning knob to select one of the following surveillance window sizes:

• NORMAL , which is 2700 feet above own aircraft and 2700 feet below ownaircraft

• ABOVE , which is 7000 feet above own aircraft and 2700 feet below own aircraft

• BELOW , which is 2700 feet above own aircraft and 7000 feet below own aircraft.

These selections are determined by the flight crew, depending on the vertical path ofthe aircraft. NORMAL would be selected during level flight. ABOVE or BELOWwould be selected during high rate climbs or descents.

The results of these two key selections will be seen on each TCAS displayseparately. In normal operation, RMU No. 1 will select these functions for the leftside TCAS display and RMU No. 2 will select for the right side TCAS display. Ifeither RMU is in the cross-side control mode (with magenta banner lines) that RMUwill control the cross-side display, just as all other cross-side controls.

Additionally, the digit in the banner line (TCAS DSPY 1 or 2) indicate which sideMFD display will be effected by the above described key press selections.

(6) PGE Key

Pressing the PGE key will bring up the RMU Page Menu. Full details of this pageare in SECTION 2, CHAPTER 7. Only those functions associated with TCAS arecovered here.

(a) ATC/TCAS Line Select Key

Pressing the ATC/TCAS line select key will display the ATC/TCAS ControlPage. There are three function select lines and one data line on this page.

1 INTRUDER ALTITUDE: (function select)

Pressing either of the adjacent line select keys will select FL. Press thiskey again will select REL. If no keys are pressed, the intruder altitude willrevert to REL after approximately 20 seconds. For a description of theseselections, refer to paragraph 3.C.(4).



2 TA DISPLAY: (function select)

a AUTO - Traffic targets displayed only when a TA or RA targetcondition exists.

b MANUAL - All traffic targets are displayed within the viewing airspace.

3 FLIGHT ID: (function select)

On this line, the Flight ID number may be entered, changed or cleared.

NOTE: The ATC FLIGHT ID function is designed foruse by the airlines. In the Embraer 145, thisfunction should be enabled. In the event thatit is inadvertently disabled, the flight ID displayline at the bottom of the transponder windowon the main page will be absent.

Press the line select key next to the FLIGHT ID data line to place anunderline cursor beneath the first character in the ID. The large tune knobmoves the underline cursor from character to character. The small tuneknob inputs the alphabetic or numeric character. Clockwise rotation of theknob will increment through the alphabet, then numerically from 0 to 9, andthen a blank. Counterclockwise rotation decrements the selectedcharacter. The blank is only available for the rightmost character.

4 FLIGHT LEVEL: (data)

This line displays the uncorrected altitude of own aircraft as is beingreported by the active transponder. The number indicates the air datacomputer from which the active transponder is obtaining its altitudeinformation. The uncorrected altitude in this example is 22,500 feet abovemean sea level, as referenced to a barometric standard of 29.92 inches ofmercury, or 1013.2 hectopascals. COM option straps and the MADCreversionary switch position determines the on-side/off-side configuration.

(b) MAINTENANCE Line Select Key (On ground ONLY)

Pressing the MAINTENANCE line select key and then the RMU SETUP lineselect key will display the RMU SETUP page. On this page, ATC FLIGHT IDmay be disabled, and therefore not transmitted in the Mode S replies. WhenATC FLIGHT ID is disabled, the FLIGHT ID legend is not shown on theATC/TCAS Control Page nor on the Main Operating page.



G. TCAS Mode Annunciations

TCAS mode annunciations are displayed on the PFD and on the MFD, when the TCASdisplay is active.

(1) On the PFD, the mode annunciation is above and to the left of the vertical speeddisplay. The annunciations are as follows:

• TCAS OFF - TCAS is not active

• TCAS TEST - TCAS TEST is active

• TCAS FAIL - TCAS Computer Unit has detected a fault

• TA ONLY - Traffic Advisory ONLY mode is active

• TCAS - TCAS is selected for display, and none of the above annunciations areactive

• RA FAIL - TCAS Computer Unit is incapable of giving Resolution Advisories

• TCAS INOP - TCAS is not enabled, but is strapped.

(2) On the MFD the annunciation is in the upper left-hand corner of the zoom window.The annunciations are as follows:

• TCAS OFF - TCAS is not active

• TCAS TEST - TCAS TEST is active

• TCAS FAIL - TCAS Computer Unit has detected a fault

• TA ONLY - Traffic Advisory ONLY mode is active

• TCAS - TCAS is selected for display, and none of the above annunciations areactive.

In TCAS AUTO the Traffic Advisory Display "Pop-up" mode is active.



3. Operation

The TCAS computer unit (CU) interrogates Mode A/C/S transponders on aircraft in the vicinityand listens for the transponder replies. By computer analysis of these replies, the CUdetermines which aircraft represent potential collision threats and provides appropriate displayindications (or advisories) to the flight crew to ensure vertical separation. Vertical separation isbased upon predictions from own aircraft altitude data and intruder altitude data supplied to theCU. The appropriate maneuver is one that ensures adequate vertical separation while causingthe least deviation of the TCAS equipped aircraft from its current vertical path.

If the threat aircraft is itself equipped with TCAS II, a coordination procedure via the air-to-airMode S data link is performed before displaying the advisory to the flight crew. This procedureassures that the advisories in each aircraft are compatible.

TCAS is designed to act as a backup to the air traffic control (ATC) system and the see andavoid concept. TCAS has a surveillance volume defined by a minimum horizontal radius of 14nautical miles and a minimum vertical range of ±12,700 feet. TCAS continually interrogatestransponders in that airspace, processes their replies, and tracks their flightpaths. Flightpathsthat are predicted to penetrate a collision area surrounding the TCAS aircraft are annunciatedboth aurally and visually. The physical dimensions of the collision area are time-based and varyas a function of horizontal and vertical closure speeds (Range Rate and Altitude Rate) andhorizontal and vertical distances (Range and Altitude) between the TCAS aircraft and theintruder aircraft.



A. TCAS Computer Unit Interface

See Figure 2-7-5.

The interface connections into and out of the TCAS computer unit are the following:

• The RSB connections are not shown on this diagram. They are shown in SECTION2.6.

• Command data into and reply data out of the TCAS computer unit is via ARINC 429links with the diversity Mode S transponders in the COM units.

NOTE: Own aircraft altitude is supplied by the air data computer via ARINC 429. Itis received by the transponder in the COM unit, which then sends the air datato the TCAS CU via the ARINC 429 link described above.

• The mutual suppression bus connects all of the transponders, DME modules, and theCU together so as to prevent them from interfering with each other.

• Weight-On-Wheels (WOW) is at ground when the aircraft is on the ground, and is usedto determine the start and end of a flight leg for fault recording.

• Input power is 115 VAC.

• Landing gear down is used in conjunction with WOW for some maintenance tests,which are further described in SYSTEM TEST AND FAULT ISOLATION, SECTION 7.

• Radio altitude and radio altitude valid are input to the CU for inhibiting certain advisorycommands below 1000 feet above ground level (AGL) and again below 400 feet AGL.

• The 8 ohm analog audio output is fed to the dedicated TCAS Speaker in the cockpit.

• ARINC 429 display data No. 1 is fed to the pilot’s IC-600 IAC where it is further sent tothe EDS for display.

• The rack-mounted cooling fan operates on 115 V ac, which comes from the CU at alltimes that power is applied.

• Aural Advisory discretes are sent to the aircraft audio system, where it is used togenerate significant audio tones.

• ARINC 429 display data No. 2 is fed to the copilot’s IC-600 IAC where it is further sentto the EDS for display.

• The 600 ohm analog audio output is fed to the aircraft Aural Warning Unit.



B. TCAS ARINC 429 Output Data

Refer to Tables 2-7-3 thru 2-7-5.

The TCAS Computer transmits ARINC 429 high speed output data to both IACs as listed inthe following table. The data is separated in the IACs for the PFDs (RA) or MFDs (TA).

Table 2-7-3. RT-910 TCAS Computer ARINC 429 Output Data Table

Parameter/Signal Name Label Rate

Control Panel Set 013

Altitude Select 015

TCAS Mode/Sens 016

Intruder Range * 130 2-3 Hz

Intruder Altitude * 131 2-3 Hz

Intruder Bearing * 132 2-3 Hz

Own Aircraft Altitude 203 2 Hz

Vertical RA 270 2-3 Hz

Horizontal RA 271 2-3 Hz

Select TCAS Sensitivity 274 2-3 Hz

Maintenance 350

STX 356

Data Characters 356

EOT 356

RTS/ETX 357

Equipment ID 377

NOTE: * Labels 130, 131 and 132 are repeated for each successive intruder fordisplay.



The TCAS Computer also transmits ARINC 429 high speed output data to the Mode STransponder as listed in Table 2-7-4. This data is for coordination with other TCAS IIequipped aircraft.

Table 2-7-4. RT-910 TCAS Computer-To-Mode S Transponder Data Table


Control Panel Set 013

Altitude Select 015

TCAS Mode/Sens 016

Own Aircraft Altitude 203 2 Hz

Vertical RA 270 2-3 Hz

Horizontal RA 271 2-3 Hz

272

273


275

276

277

Maintenance 350

The Mode S Transponder transmits ARINC 429 high speed output data to theTCAS Computer as listed in Table 2-7-5. This data is for coordination with other TCAS IIequipped aircraft.

Table 2-7-5. RCZ-851E Communications Unit-To-TCAS Computer Data Table


273


275



RMP LBP RMP RBP

RT-300 RADIO ALTIMETER R/T

IC-600 IAC NO. 2


RT-300 RADIO ALTIMETER R/T

IC-600 IAC NO. 1

RT-910 MOUNTING RACK FAN

RT-910 TCAS COMPUTER UNIT

15J

15K

12

5

X

N

Y

29

30

5

9

2H

2J

2K

5K

1

7

2F

2G

7C

7D

14F

14G

13F

7E

7J

6K

3A

3B

3C

1F

2A

7G

7H

X

N

Y

29

30

C20J1

C190J2B

(H)

(L)

(H)

(L)

(H)

(C)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

FROM 115 V AC BUS NO. 1

115 V AC RETURN


TO ALL L-BAND EQUIPMENT

AUXILIARYOUTPUT

VALID

(H)

(L)

ARINC 429 #1CO-ORD OUTPUT

MUTUAL SUPPRESSION

RADIOALTITUDE

+28 V = VALID

W.O.W.

115 V ACINPUT POWER

115 V ACRACK FAN POWER

ARINC 429 #1CO-ORD INPUT

LDG GEAR DISCRETE

RADIOALTITUDE

+28 = VALID

AUXILIARY OUTPUT

VALID

193J1193J1

20J1

190J2B

115 V AC

NO. 1

NO. 2

TCAS SPEAKER

55

87

90

104

143J1

(H)

(L)

(H)

(L)

NOTE:


55

87

90

104

C143J1

(H)

(L)

(H)

(L)



14H

14J

14A

14B

(H)

(L)

(H)

(L)

ARINC 429 #2CO-ORD OUTPUT

ARINC 429 #2CO-ORD INPUT

MUT SUP94

ARINC 429TX CO-ORD

OUTPUTARINC 429

XT CO-ORDINPUT

94MUT SUP

TRAFFIC

ARINC 429TA/RA #2

DISPLAY DATA

8 OHMAUDIO OUTPUT

ALTITUDE DATA FROM MADC TOCOMM UNIT VIA ARINC 429.

14C

13E

1K

AUDIO SYSTEM

CORRECTIVEAURAL

ADVISORIES

PREVENTIVE

CHASSIS GROUND

SIGNAL GROUND8

11

GROUND = GEAR DOWN

FROM GEAR HANDLE SWITCH

TA/RADISPLAY DATA

TA/RA #1DISPLAY DATA

AURAL WARNING UNIT

3F

3G

(H)

(L)

(H)

(L)600 OHM

AUDIO OUTPUT

TA/RADISPLAY DATA

PROGRAM COMMON

4GRA SELF-TEST MONITOR INHIBIT

ENABLES TA & RAOUTPUT TO IACS

12CEFIS OUTPUT ENABLE

AD-49870-R1@

Figure 2-7-5. TCAS Computer Unit Interface Diagram



C. TCAS Displays

TCAS operational displays are divided into two distinct advisories. These are trafficadvisories (TAs) and resolution advisories (RAs).

(1) Traffic Advisories (See Figures 2-7-6 and 2-7-7.)

TAs are announced aurally and are shown on the multifunction displays (MFDs)when selected by the flight crew. The flight crew uses this information only as an aidto visually locate the intruder in order to avoid a conflict.

The TA display includes the range and bearing of the intruder relative to the TCASaircraft. If the intruder is equipped with altitude reporting capability, intruder altitudeis displayed either as altitude relative to the TCAS aircraft or as uncorrected altitudeof the intruder.

The TCAS TA displays use color-coded symbols and data tags to map air traffic andlocal threat aircraft on the MFD.

Four traffic symbols are used: solid circle, solid square, solid diamond, and hollowdiamond. A different color is assigned to each symbol type, as listed in Table 2-7-6.

Table 2-7-6. MFD/TCAS Symbology

Graphic Symbol Color Display Function

Solid Square Red Resolution Advisory (RA)

Solid Circle Amber Traffic Advisory (TA)

Solid Diamond Blue Proximate Traffic

Hollow Diamond Blue Other Traffic



TCAS TESTABVFL

RA NO BRGTA NO BRG

+20

-10

-05

-12

MAP/PLANDISPLAY AREA


00

AD-51321@

TCAS MODEANNUNCIATIONS

(TCAS TEST-AMBER)TCAS FAIL-AMBER)TCAS OFF-WHITE)

TA ONLY-WHITE)TCAS-WHITE)

RA SYMBOL(RED)

OA SYMBOL(CYAN)

AIRPLANE SYMBOL(WHITE)

2 NM RANGERING (WHITE)

TA SYMBOL(AMBER)

PA SYMBOL(CYAN)

TCAS ZOOMWINDOW(DIM-WHITE)

TCAS RANGESCALE (WHITE)

OFF-SCALEINTRUDER

6

Figure 2-7-6. TCAS MFD Symbology (Sheet 1)



TCAS TESTABVFL

RA NO BRGTA NO BRG

MAP/PLANDISPLAY AREA


AD-51320@

25

-12-05

00

00

ABOVE/BELOWDISPLAY MODE

(WHITE)

ABSOLUTEALTITUDE

ANNUNCIATION(WHITE

TCAS SCALEHALF-RANGEARC (WHITE)

RELATIVEALTITUDEDISPLAY

VERTICALSENSEARROW

TCAS RANGEDIGITAL READOUT(WHITE)

TA NO BEARINGDISPLAY (AMBER)

RA NO BEARINGDISPLAY (RED)

Figure 2-7-7. TCAS MFD Symbology (Sheet 2)



(2) Resolution Advisories (See Figure 2-7-8.)

RAs are also announced aurally and are shown on the primary flight displays (PFDs)as part of the Vertical Speed Indicator (VSI), whenever the TCAS computer detectsan imminent collision. The flight crew follows these commands unless they havecertain knowledge that doing so will jeopardize the continued safety of the flight.

(a) Traffic Advisory Colors

1 Red

Represents an immediate threat to a TCAS-equipped aircraft. Promptaction is required to avoid an intruder. This color is only used inconjunction with an RA.

2 Amber

Represents a moderate threat to a TCAS-equipped aircraft. A visualsearch is recommended to prepare for intruder avoidance. Amber is usedonly in conjunction with a TA.

3 Cyan

Represents proximate traffic and other traffic that the TCAS surveillancelogic has in its track file.

4 White

Used only for mode annunciations and for reference graphics, includingown aircraft position, range ring, etc.

(b) Traffic Identification

1 Resolution Advisory

Intruder aircraft entering the warning area, 15 to 35 seconds from theTCAS II collision area, are represented as a solid red square. This type oftraffic will result in an RA.

2 Traffic Advisory

Intruder aircraft entering the caution area, 20 to 48 seconds from the TCASII collision area, are represented as a solid amber circle. This type oftraffic will result in a TA.

3 Proximate Traffic

Aircraft within display range and within the selected vertical window arerepresented as a solid cyan diamond. Proximate traffic is shown toimprove situational awareness in the event of a potential conflict withhigher priority RA or TA aircraft.



32

1

0

3

1

2

1000

TCASTEST

AD-51322@

TCAS MODEANNUNCIATIONS

(TCAS TEST-AMBER)(TCAS FAIL-AMBER)(TCAS OFF-WHITE)

(RA FAIL-AMBER/RED)(TCAS INPO-WHITE)

TCAS "FLT-TO" ZONEBAND (GREEN)

VERTICAL SPEEDIN RED BAND(POINTER-RED)DIGITS-RED)NOT IN RED BAND(POINTER-WHITE)(DIGITS-WHITE)

TCAS RA BAND(RED)

Figure 2-7-8. TCAS PFD Symbology

4 Other Traffic

Any transponder-replying traffic that is not classified as an intruder orproximate traffic, is within the display range, and is within the selectedvertical window are represented as hollow cyan diamonds (only in viewwhen no RA or TA is in progress). The predicted flight paths of proximatetraffic and other traffic do not penetrate the collision area of the aircraft.

(c) Data Tags

Two styles of data tags may be displayed on the MFD. The style displayed isdetermined by the RMU selection of ALT:REL or ALT:FL. Refer to paragraph3.E.(6)(a) for selection procedure.

When ALT:REL is selected, a data tag , made up of a 2-digit number and aplus (+) or a minus (-) sign, appears either above or below the intruder aircraftsymbol and represents the relative altitude of the intruder, in hundreds of feet,as referenced to the TCAS equipped aircraft (+ means that the intruder isabove the TCAS equipped aircraft, and - means that the intruder is below).The data tag may also include a vertical arrow. If the arrow is pointingupward, it means that the intruder aircraft is climbing at a rate greater than500 feet per minute (fpm) and if the arrow is pointing down, the intruder isdescending at a rate greater than 500 fpm. The data tag appears in the samecolor as the advisory.

When ALT:FL is selected, the data tag is made up of a 3-digit number, whichrepresents the intruders reported uncorrected altitude in hundreds of feetabove mean sea level.



D. TCAS Display Logic

(1) Traffic Advisory (TA) Logic

TCAS monitors a time-based dimension of a caution area that extends 20-48seconds from the time the intruder is predicted to enter the TCAS aircraft’s collisionarea. Should an intruder aircraft enter this caution area, traffic information in theform of a traffic advisory (TA) is issued by the TCAS computer. This information isconveyed to the flight crew via the MFD as described in paragraph 3.B.(1)(a).

(2) Resolution Advisory (RA) Logic

TCAS also monitors a time-based dimension of a warning area that extends 15-35seconds from the time the intruder is predicted to enter the TCAS aircraft’s collisionarea. Should an intruder enter this warning area, an escape strategy in the form of aresolution advisory (RA) is issued by the TCAS computer. This information isconveyed to the flight crew via the VSI as described in paragraph 3.B.(1)(b).

The appropriate intruder TA display symbol on the MFD changes to a red square.

E. TCAS Preflight Test

Basic TCAS operating procedures on the ground include preflight test, TCAS activationbefore takeoff, and TCAS deactivation after landing. In-flight procedures are contained inthe pilot’s manual.

The Pilot-Activated Self-Test (PAST) feature is a convenient way to test the TCAS systembefore takeoff.

NOTE: PAST can be initiated at any time, on the ground or in flight, either active orin standby. If while in flight, TAs or RAs occur during self-test, the test willabort and the advisories will be processed and displayed as appropriate.

(1) Activate TCAS

On the RMU, select the transponder which will cause the banner line to displayATC/TCAS and place it in an active TCAS mode.

(2) Activate TCAS Display

(a) From the MAIN 1/2 MFD menu, select MFD SETUP.

(b) From the MFD SETUP menu, toggle TRAFFIC ON.



(3) Self-test

(a) Press a line select key to place the cursor in the bottom left window.

(b) Press and hold the TST key.

(c) Observe the following:

1 Aural annunciation TCAS TEST is heard

2 Both VSIs display red and green resolution advisory overlay indicating"don’t descend, don’t climb > 2000 ft/min" advisory.

(d) The on-side MFD will display the following test pattern:

• An RA at 3 o’clock, 2 NM, 200 feet above, in level flight

• A TA at 9 o’clock, 2 NM, 300 feet below, climbing

• A PA at 3.6 NM, 33 degrees right of the aircraft heading (approximately 1o’clock), 1100 feet below, descending

• A non-threat intruder at 3.625 NM, 33.75 degrees left of aircraft heading(approximately 11 o’clock), 2000 feet above, in level flight

• TCAS TEST annunciation on PFD and on MFD.

(e) After 8 seconds, aural annunciation TCAS PASS is heard, and the testpatterns are replaced by normal VSI and MFD displays.

(f) If self-test fails, TCAS TEST FAIL is announced aurally and TCAS FAILappears on the PFD and on the MFD.



4. Fault Monitoring

The TCAS II is programmed to perform automatic self-tests at power-on and during operation.The tests performed include power on self-test, continuous self-test and commanded self-test.

A. Fault Detection

The CU detects system faults and reports them on its front panel lamp display. Its flightleg memory stores system status and fault information for 10 consecutive flight legs. Aflight leg is the interval between weight-off-wheels and weight-on-wheels during whichTCAS is operative. By recalling the stored data, ground maintenance personnel canevaluate in-flight performance on the ground and fault isolate a current or previous failureto a specific LRU or LRU interface.

B. Fault Indications

Fault indications are presented on both the PFDs and the MFDs. In both cases, themessage TCAS FAIL will be displayed as described in paragraph 2.G.(2). The trafficadvisory and resolution advisory displays will be removed.





2.8 FLIGHT MANAGEMENT SYSTEM (FMS) - Optional . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-1


A. NZ-2000 Navigation Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-2

B. IM-803 Configuration Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-4

C. CD-810 Control Display Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-5

D. DL-900 Data Loader . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-10

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-11

A. FMS Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-13

B. FMS ARINC 429 Input/Output Data . . . . . . . . . . . . . . . . . . . . . . . . 2-8-14



B. MFD Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-21




FIGURE/TITLE PAGE


Figure 2-8-1. NZ-2000 Navigation Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-2

Figure 2-8-2. IM-803 Configuration Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-4

Figure 2-8-3. CD-810 Control Display Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-5

Figure 2-8-4. DL-900 Data Loader (Access Door Open) . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-10

Figure 2-8-5. FMS Interface Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-17/18

Figure 2-8-6. PFD FMS Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-20

Figure 2-8-7. MFD FMS Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-22

List of Tables

TABLE/TITLE PAGE

Table 2-8-1. NZ-2000 Navigation Computer Leading Particulars . . . . . . . . . . . . . . . . . . . . 2-8-3

Table 2-8-2. IM-803 Configuration Module Leading Particulars . . . . . . . . . . . . . . . . . . . . 2-8-4

Table 2-8-3. CD-810 Control Display Unit Leading Particulars . . . . . . . . . . . . . . . . . . . . . 2-8-6

Table 2-8-4. DL-900 Data Loader Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-10

Table 2-8-5. FMS Navigation Computer ARINC 429 Output Data Table . . . . . . . . . . . . . . 2-8-14

Table 2-8-6. FMS Navigation Computer Unused ARINC 429 Output Data

Table . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-15

Table 2-8-7. IAC to Navigation Computer ARINC 429 Input Data Table . . . . . . . . . . . . . . 2-8-16

Table 2-8-8. PFD Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-19

Table 2-8-9. MFD Failure Indications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8-21



SECTION 2.8FLIGHT MANAGEMENT SYSTEM (FMS) - Optional

1. General

The PRIMUS® 1000 Integrated Avionics System installed in the Embraer 145 aircraft includesan optional flight management system (FMS).

The primary function of the FMS is to provide high accuracy lateral and vertical navigation fromany point in the world, to any other point in the world. The FMS provides many additionalfunctions such as; remote radio tuning, flight plan building and storage, waypoint creation andstorage, and information on NAVAIDs and earth reference points, such as airports, namedintersections, VOR/DMEs, runways, and routes.

To accomplish this function, the FMS navigation computer must calculate and maintain anaccurate position. Although the FMS interfaces with a variety of short range and long rangesensors, the sensors themselves are not part of the FMS. These sensors include VOR, DME,AHRS, and GPS. Each sensor has individual characteristics which allow sensors to complementeach other. For example, the AHRS has very good short-term accuracy, but has poor long-termaccuracy, while DME has good long-term accuracy but has poor short term accuracy. By usinga weighted average of the sensor inputs, the navigation computer can generate an FMS positionwhich is as accurate as any single sensor under any given condition. Control of these sensorsis managed through the Control Display Unit (CDU) which is described later.

The lateral navigation function of the FMS may be considered an area NAV system (RNAV). Itsfundamental purpose is to provide navigation information relative to selected geographic points.Navigation management allows the pilot to define a route from aircraft present position to anypoint in the world. The FMS will output advisory information and steering commands to allowthe pilot or the autopilot to guide the aircraft along the desired route. Routes are defined fromaircraft present position to a destination waypoint, via a great circle route or a series of greatcircle legs, defined by intermediate waypoints.

The FMS has in its memory two databases. A navigation database that contains data onNAVAIDs, airports, and airways and is updated every 28 days and a custom database that thepilot uses to create and store flight plans and waypoints. The custom database is not updatedon any scheduled basis.

The optional installation consists of the following LRUs:

• One NZ-2000 Navigation Computer

• One CD-810 Control Display Units (CDUs)

• One IM-803 Configuration Module

• One DL-900 Data Loader.

For full operational information on the FMS, refer to Honeywell Pub. No. 28-1146-043,FMZ-Series Flight Management System Pilot’s Manual.




A. NZ-2000 Navigation Computer

Figure 2-8-1 shows a graphical view of the NZ-2000 Navigation Computer. Table 2-8-1provides items and specifications that are particular to the computer.

AD-39956@

Figure 2-8-1. NZ-2000 Navigation Computer



Table 2-8-1. NZ-2000 Navigation Computer Leading Particulars

Item Specification


• Length . . . . . . . . . . . . . . . . . 17.25 in. (438.2 mm)

• Width . . . . . . . . . . . . . . . . . . 2.28 in. (57.9 mm)

• Height . . . . . . . . . . . . . . . . . . 7.62 in. (193.5 mm)

Weight . . . . . . . . . . . . . . . . . . . . 7.94 lb (3.17 kg)

Power Requirements:

• Continuous . . . . . . . . . . . . . . 28 V dc, 48 W (nom)

• Surge . . . . . . . . . . . . . . . . . . 28 V dc, 55 W (max)

User Replaceable Parts . . . . . . . . . None

Mating Connector:

• J1 . . . . . . . . . . . . . . . . . . . . . Cannon Part No. DPX2-A106P-A106P-33B-0068

Mounting . . . . . . . . . . . . . . . . . . . Tray, HPN 7020157-901

The navigation computer receives its FMS command data from:

• CD-810 Control Display Unit (CDU).

The navigation computer receives its FMS input data from:

• AH-800 AHRU and AZ-850 MADC via ARINC 429

• RNZ-851 NAV Unit via CSDB

• DL-900 Data Loader via RS-422

• Global Navigation System Sensor Unit (GNSSU) via ARINC 429

• Both IC-600 IACs.

The navigation computer contains the necessary power supplies, electronics, and databasememory to receive and process sensor input information, to provide highly-accuratepresent-position information to the flight crew. Additionally, the navigation computer hasthe ability to remotely tune all the radios on the aircraft, as well as provide a means for theflight crew to create and store waypoints and flight plans.

Radio tuning command data is via a CSDB link between the navigation computer and theRMUs. The data is then sent to the remote units via RSB.

The data transfer link between the CDU and the NZ-2000 is over an RS-422 private lineinterface. The FMS does not display navigation maps on the CDU, however, the FMS isthe source of map data for EDS displays.



B. IM-803 Configuration Module

Figure 2-8-2 shows a graphical view of the IM-803 Configuration Module. Table 2-8-2provides items and specifications that are particular to the module.

AD-42378@

Figure 2-8-2. IM-803 Configuration Module

Table 2-8-2. IM-803 Configuration Module Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.98 in. (24.98 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.52 in. (114.8 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.83 in. (173.5 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.6 lb (0.27 kg)

Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Supplied by NZ-2000

User Replaceable Parts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . None

Mating Connector . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cannon DEMF-9S

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Four Mounting Screws

The Navigation Computer (NZ) has a number of options which may be selected by theinstaller by means of configuration straps. These straps are located on a printed wiringassembly which is contained within the IM-803 Configuration Module, and are aircraftunique. Strap programming procedures are covered in SECTION 4 - MAINTENANCEPRACTICES in this manual.



C. CD-810 Control Display Unit

Figure 2-8-3 shows a graphical view of the CD-810 Control Display Unit. Table 2-8-3provides items and specifications that are particular to the unit.

AD-11942-R1@

NEXT FPL PROG DIRPREVNAVPERF

PHOTOSENSOR ANNUNCIATORS

PHOTOSENSOR

CRTDISPLAY

LEFTLINE

SELECTKEYS

SCRATCHPAD

MODE KEYS

ALPHA-NUMERICKEYS

FUNCTIONKEYS

RIGHTLINESELECTKEYS

FUNCTIONKEYS

MODE KEYS

BRIGHTNESSCONTROL

DSPLY DR DGRAD MSG OFFSET APRCH

BRT

A

G

M

B

H

N

C

I

O

D

J

P

E

K

Q

F

L

R

S

X

T

Y

U

Z

V W

DEL CLR

1 2 3

4 5 6

7 8 9

0

/

ACTIVE FLT PLAN

MLF

SLC

KSLC

KSLCARM ALTN

1/5DEST

61Ø16Ø/492Ø

456Ø

ARRIVAL

02Ø˚ 93.ØNM

348 4.ØNM

Figure 2-8-3. CD-810 Control Display Unit



Table 2-8-3. CD-810 Control Display Unit Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . 7.5 in. (190.5 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . 5.75 in. (146.1 mm)

• Length (from rear of bezel) . . . . . . . . . . 10.0 in. (254.0 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.7 lb (5.76 kg)

Power Requirements:

• Primary . . . . . . . . . . . . . . . . . . . . . . . . 28 V dc, 46 W (max)

• Lighting . . . . . . . . . . . . . . . . . . . . . . . . 28 V dc, 13 W (max)


• Knob, Brightness Control . . . . . . . . . . . HPN 7008508

• Setscrew, 4-40 X 1/8-inch . . . . . . . . . . . MS51021-9, HPN 0455-128

Mating Connector:

• J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS3126F22-55SX

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . Standard Dzus Rail

The CDU consists of a keyboard, a color cathode ray tube (CRT) display, and theelectronics required to communicate with the navigation computer.

The CDU provides the primary means for pilot input to the system. It also provides outputdisplay for the navigation computer. The CDU utilizes a full alphanumeric keyboard, aswell as decimal, dash, and slash. Four line selection keys are provided on each side of theCRT. Seven function keys are provided to allow direct access to specific display pages.Annunciators are located in the top of the bezel to advise the pilot of the system’s status.

The CRT in the CDU has nine lines of text, 24 characters long. The top line of the CDUdisplay is dedicated as a title line and the bottom line is used as a scratchpad and todisplay messages. A manual dimming knob is used for long-term dimming adjustments,while ambient light sensors are used for short-term display brightness adjustments undervarying cloud/sunlight conditions.

The use of colors on the CDU is designed to highlight important information. Colorassignments are coordinated as much as possible with the electronic display system(EDS).



(1) CDU Color Assignments

The following general rules are used for assigning colors:

• Vertical - Cyan (blue)

• Lateral - Green

• FROM Waypoint - Yellow

• TO Waypoint - Magenta

• Prompts and Titles - White

• Flight Plan Names - Orange

• Index Selections - Green.

(2) Scratchpad

The bottom line of the CRT is the scratchpad, which provides a working area for thepilot to enter data and/or verify the data before line selecting it to the desiredposition on the display. The scratchpad also allows advisory and alerting messagesto be displayed to the pilot.

Alphanumeric entries are made to the scratchpad via the keyboard. As each key isdepressed, that character is displayed in the scratchpad. Information in thescratchpad does not affect the FMS until it is line key selected to a line on thedisplay.

(3) Line Select Keys

There are four line select keys on each side of the CRT display. Data may beselected from the scratchpad to a line, or from a line to the scratchpad through theuse of these keys.

(4) Function Keys

(a) Clear Key (CLR)

This key has the following functions:

• When a message is present in the scratchpad, depressing the CLR keyclears that message

• When an entry beginning with an asterisk (*) or (#) is in the scratchpad,depressing the key removes the entire entry

• When an alphanumeric entry is in the scratchpad, one character at a timeis cleared from the scratchpad (from right to left) for each time the buttonis depressed

• When an alphanumeric entry is in the scratchpad and the CLR key is helddown, the first character is cleared. After approximately a 1/2 second haspassed, characters will be cleared for as long as the CLR key is helddown.



(b) Delete Key (DEL)

The delete key has one function. When there is no message in the scratchpadand the DEL key is depressed, *DELETE* will appear in the scratchpad. Thismay now be line selected to delete waypoints and other data displayed in theCRT data fields.

When there is a message displayed, the delete operation is inhibited.*DELETE* is also used to return to default values after entries have beenmade.

(c) Performance Key (PERF)

Depressing the PERF function key enables the pilot to access the performanceindex. The pilot may select any of the PERF submodes by depressing theappropriate line select key.

(d) Navigation Key (NAV)

Depressing the NAV function key enables the pilot to access the NAV indexpage. The pilot may select any of the NAV submodes by pressing theappropriate line select key.

(e) Flight Plan Key (FPL)

Pressing the FPL key displays the first page of the active flight plan. If thereis no flight plan currently entered, the pilot may manually enter a flight plan,load a flight plan from a diskette, or select a stored flight plan.

(f) Progress Key (PROG)

Depressing the PROG key displays the first of the progress pages. The firstprogress page displays the ETE, distance to, and fuel projections for the TOwaypoint and destination; the current NAV mode; the number of long rangeNAVs used and the NAVAIDs that are currently tuned for radio updating.

(g) Direct TO/Intercept Key (DIR)

Depressing the DIR function key displays the active flight plan with theDIRECT, HOLD, and INTERCEPT prompts.



(5) Annunciators

The six annunciators located at the top of the CDU operate independently from theCRT and keyboard. Illumination of the annunciators is initiated by the navigationcomputer. Except for the display (DSPLY) annunciation, an annunciation is alsoprovided on the PFD. The two colors used for the annunciations are white andamber. White indicates an advisory annunciation, while amber indicates an alertingannunciation.

(6) Brightness Control

Brightness control is provided for the CDU CRT display in order to maintainreadability under dim light, as well as direct sunlight. This is accomplished in twoways:

• Manually by the brightness knob

• Automatically by photosensors.



D. DL-900 Data Loader

Figure 2-8-4 shows a graphical view of the DL-900 Data Loader. Table 2-8-4 providesitems and specifications that are particular to the data loader.

AD-29793@

Figure 2-8-4. DL-900 Data Loader (Access Door Open)

Table 2-8-4. DL-900 Data Loader Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.24 in. (56.90 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.75 in. (146.05 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.00 in. (203.20 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.0 lb (1.36 kg)

Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . 28 V dc, 9 W (max)

User Replaceable Parts . . . . . . . . . . . . . . . . . . . . . . . . None

Mating Connector:

• J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS3126F16-26S




The DL-900 Data Loader is used to transfer navigation related data to the FMS navigationcomputer. The DL-900 uses 3.5-inch diskettes and has an RS-422 interface with thenavigation computer.

(1) Navigation Database Loading

The DL-900 Data Loader provides transfer of data derived from the Jeppesendatabase. This data includes NAVAIDs, waypoints, airports, runways, procedures,and jet routes. The database is updated every 28 days.

(2) Flight Plan Loading

The DL-900 Data Loader also has the capability of interfacing with a Lockheed JetPlan Computer or equivalent. This allows the pilot the option of loading a flight planfrom a diskette.

(3) GNSSU Software Loading

The DL-900 Data Loader also has the capability of interfacing with the optionalGlobal Navigation System Sensor Unit (GNSSU) for software updates.

3. Operation

Full information on the operation of the FMS, including the CDU and Data Loader, are inHoneywell Pub. No. 28-1146-043, FMZ-Series Flight Management System Pilot’s Manual. Dueto the complexity of the operation, it is not duplicated here.

The primary FMS task is to navigate the aircraft along a predefined flight plan. To do this, theFMS receives navigation data from various sensors onboard the aircraft and chooses thesensors providing the most accurate aircraft position.

The sensor selections of the FMS in the Embraer 145, listed in priority order are as follows:

• GPS with RAIM

• GPS without RAIM

– DME/DME/GPS

– VOR/DME/GPS

• DME/DME

• VOR/DME.

The FMS will operate in the highest priority mode supported by the available sensors. Indetermining the navigation mode, the FMS evaluates the accuracy of each sensor and usesthose that will yield the best FMS position.



The FMS also provides the ability to automatically tune the aircraft’s VOR and DME receivers.Calculation of aircraft present position from VOR/DME information requires input of bearing anddistance data, as well as knowledge of the station’s coordinates in latitude and longitude. Thenavigation database is periodically used by the navigation computer to find the coordinates andfrequency of the high and low altitude VORTAC and VOR/DME stations in the vicinity of theaircraft. When the desired stations are chosen, the frequency is output to the navigationreceivers through the RMU. Automatic receiver tuning is operationally transparent to the pilot,other than periodic changes in the receiver’s displayed frequency and the RMI pointer.Provision is included for remote tuning of the receivers via the CDU, through the PRIMUS® IIRadio Management Unit (RMU). For remote tuning via the CDU, the pilot can choose to enterthe station identifier or frequency.

An important part of the navigation computer is the non-volatile memory area or database whichcontains information on NAVAIDs, airports, and airways. The database is integral to thenavigation computer to allow quick access of the stored information. The database is updatedevery 28 days for accuracy.

The database area is also used to store pilot-defined waypoints and flight plans. This"CUSTOM" database is not updated every 28 days. This data stays in memory until changed bythe pilot.

The navigation database contains the following information:

• VORs - Worldwide

• ILS/MLS - Worldwide

• NDBs - Worldwide

• Airports - Worldwide (that meet certain criteria)

• Runways - Worldwide

• Airways - Worldwide (both high and low altitude)

• SIDs/STARs - Worldwide (that are published)

• Approach Waypoints - Worldwide

• Named Intersections - Worldwide.

The navigation computer provides a lateral steering command to the flight directors in the IACs.This lateral steering command can be manually flown by the pilot or sent to the autopilot forautomatic flight path steering. The lateral steering command is proportional to the calculateddistance and angle deviation from the desired lateral track.



A. FMS Interface Diagram

See Figure 2-8-5.

NOTE: All the data buses and their characteristics are covered in SECTION 1 of thismanual.

Command data for the FMS NAV Computer originates in the CDU. The data is sent to thenavigation computer (NZ). Reply data and display data for the CDU are sent from the NZto the CDU via another RS-422 data bus. There are five shielded-twisted pairs included inthis RS-422 data bus structure.

The data loader communicates with the navigation computer via an RS-422 data bus.

The navigation computer sends ARINC 429 data to the IACs for display of map data on thePFD and MFD for long-range lateral navigation guidance. This data is also sent to theGround Proximity Warning System (GPWS).

The navigation computer receives data from each IAC on separate ARINC 429 databuses.This data includes such information as the designator latitude/longitude, and fuel flow.

The FMS NAV Computer sends command data to the VOR and DME receivers through theRMUs, and reply VOR and DME data is sent back via CSDB.

The global navigation system sensor unit (GNSSU) accepts position data from the FMS viaan ARINC 429 bus. This data is used in the initialization process and is reported back onan ARINC 429 bus as FMS data. Once the GNSSU enters the navigation mode, the datais reported as GPS data and the FMS then uses the data for its navigation functions.

Altitude, true airspeed, and vertical velocity from the MADCs are sent to the FMS viaARINC 429.

Heading data is sent from the AHRS to the FMS via ARINC 429.

Input/Output (I/O) to the Automatic Flight Information System (AFIS) is also via ARINC 429.

Other connections include power and ground which have been discussed previously.

Weight On Wheels (WOW) is used by the navigation computer to determine which sensorsto use for position updates when on the ground.

The navigation computer has a number of options which are selected by the installer bymeans of configuration straps. These straps are located on a printed wiring assemblywhich is contained within the IM-803 Configuration Module, and are aircraft unique. Strapprogramming procedures are covered in SECTION 4, MAINTENANCE PRACTICES, of thismanual.



B. FMS ARINC 429 Input/Output Data

Refer to Tables 2-8-5, 2-8-6, and 2-8-7.

(1) Table 2-8-5 lists the ARINC 429 data that the navigation computer transmits to theIACs.

Table 2-8-5. FMS Navigation Computer ARINC 429 Output Data Table

Label Label Description

075077101114115116117121122147251252275277312315316321326327372373

WPT/OBSApproach Reference Speed BugSelected HeadingDesired TrackWaypoint BearingCrosstrack DistanceVertical DeviationRoll SteeringPitch SteeringMag-VarDistanceTime To GoTo/From, APP, XTRK, DGR, DRVNAV Submode AnnunciationGround SpeedWind SpeedWind DirectionDrift AngleLateral Scale FactorVertical Scale FactorFHDG Submode Characters 1-3FHDG Submode Characters 4-6

The following labels support map data:

074113300301302303304305306307310311

Data Record HeaderMessage ChecksumSTN Declination, Type and ClassMessage Characters (7-9)Message Characters (10-12)Message Length/type/numberMessage Characters (1-3)Message Characters (4-6)NAV/WPT/AP LatitudeNAV/WPT/AP LongitudePresent Position LatitudePresent Position Longitude



(2) Table 2-8-6 lists the ARINC 429 data that the navigation computer transmits to theIACs, but are not currently used by the IACs.

Table 2-8-6. FMS Navigation Computer Unused ARINC 429 Output Data Table

Label Label Description

125163204210213260313314320351352371

Greenwich Mean Time (GMT)Wind On NoseSystem AltitudeTrue AirspeedStatic Air TemperatureDateTrue Track AngleTrue HeadingMagnetic HeadingDistance to DestinationTime to DestinationEquipment Identification



(3) Table 2-8-7 lists the ARINC 429 data that the IACs transmit to the navigationcomputer.

Table 2-8-7. IAC to Navigation Computer ARINC 429 Input Data Table

ARINCLabel

Parameter Comments

100101102203204205206210211212213270306307320324325333371314030*030*031*031*032*032*034*034*247347*347*107TBDTBD

Selected CourseSelected HeadingSelected AltitudePressure Altitude (from disp ADC)Baro Corrected Altitude (disp ADC)Mach (disp ADC)Indicated Airspeed (disp ADC)True Airspeed (disp ADC)Total Air Temperature (disp ADC)Altitude Rate (disp ADC)Static Air Temperature (disp ADC)DGC Status (IC-600 Status)Designator LatitudeDesignator LongitudeMagnetic Heading (disp AHRS)Pitch Attitude (disp AHRS)Roll Attitude (disp AHRS)Body Normal Accel (prim AHRS)Equipment IDTrue Heading (disp AHRS)VHF COM Frequency #1VHF COM Frequency #2Beacon Transponder Code #1Beacon Transponder Code #2ADF Frequency #1ADF Frequency #2VOR/ILS Frequency #1VOR/ILS Frequency #2Total Fuel QuantityFuel Flow Engine #1Fuel Flow Engine #2Flap PositionBleed StatesVmo/Mmo

* means transmitted with SDI

Existing/GAMA 429 2nd Ed.Existing/GAMA 429 2nd Ed.Existing/GAMA 429 2nd Ed.Existing/ARINC 429-13Existing/GAMA 429 2nd Ed.Existing/ARINC 429-13Existing/ARINC 429-13Existing/GAMA 429 2nd Ed.Existing/ARINC 429-13Existing/ARINC 429-13Existing/GAMA 429 2nd Ed.Non-standard bit patternExisting/GAMA 429 2nd Ed.Existing/GAMA 429 2nd Ed.Existing/GAMA 429 2nd Ed.Existing/ARINC 429-13Existing/ARINC 429-13Existing/ARINC 429-13Existing/GAMA 429 2nd Ed.



1

14

13

26

92

52

103

104

57

56

28

62

49

M

N

S

T

P

R

U

V

W

X

i

T

S

H

G

K

J

E

86

72

45

58

29

28

65

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

CDU VALID

LOADER CONNECTED LEFTDL CONNECTED

CDU VALID

A B121J1

120J1

123J1

WOW 77

47

84

SDI NO. 1

SDI NO. 2

A B121J1

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

(H)

(L)

96

82

80

81

94

95

73

74

17

18

ARINC 429 XMTRGEN BUS NO. 1

CSDB RCVRNAV PRIMARY

CSDB RCVR

NAV SECONDARY

CSDB RCVR

DME PRIMARY

CSDB RCVRDME SECONDARY

ARINC 429 RCVRLTS NO.1

RS422 RCVRCDU DATA

RS422 XMTRCDU DATA

RS422 RCVRCDU CNTL

RS422 XMTRCDU CNTL

RS422 XMTRCDU CLK

RS422 RCVRDATA LOADER

RS422 XMTRDATA LOADER

RS422 XMTRCLOCK

RS422 XMTRDATA

RS422 RCVRDATA

RS422 XMTRCNTL

RS422 RCVRCNTL

RS422 RCVRCLK

RS422 XMTRDATA

RS422 RCVRDATA

RS422 RCVRCLK

NZ-2000 NAVIGATIONCOMPUTER (Optional)

CD-810 CDU NO. 1

DL-900DATA LOADER

NC

105

91

40SDI NO. 3 NC

61

8

48

RS232 DATA

5 VDC PWR IN

SIGNAL GND

IM-803 INSTMODULE

RS232 DATA IN

5 VDC INST MODULE PWR

INST MODULE GND

4

16

(H)

(L)

ARINC 429 XMTRGEN BUS NO. 3

49

50

44

43

(H)

(L)

ARINC 429 RCVRAHRS NO. 4

(H)

(L)

ARINC 429 RCVRAHRS NO. 5

199J1

3

1

2

(H)

(L)

ARINC 429 RCVRNO. 15

18

37

149J1

GNSSU(H)

(L)ARINC 429 FMSINPUT NO. 1

(H)

(L)

38

39

MADC NO. 1 (H)

(L)

68

69

MADC NO. 2 (H)

(L)

9J1

68

69

68 (L)

ARINC 429 RCVRMADC NO. 12

67(H)

ARINC 429N O. 4 OUT

ARINC 429N O. 4 OUT

76 (L)

ARINC 429 RCVRMADC NO. 13

90 (H)

AHRU NO. 1 (H)

(L) K13

ARINC 429N O. 3 OUT

K12

AHRU NO. 2 (H)

(L) K13

ARINC 429N O. 3 OUT

K12

1J1B

C1J1B

RNZ-851 NO. 1

RNZ-851 NO. 2

(H)

(L)

C164J1

164J1

B89

A59

(H)

(L)

A57

A44

CSDBVOR OUT

CSDBDME OUT

(H)

(L) A59

(H)

(L)

A57

A44

CSDBVOR OUT

CSDBDME OUT

B89

IAC NO. 1

IAC NO. 2

(H)

(L)

C190J2A

190J2A

(H)

(L)

(H)

(L)

(H)

(L)

39

52

21

22

21

22

39

52

ARINC 429LRN OUT

ARINC 429CH0 IN

ARINC 429LRN OUT

ARINC 429CH0 IN

ARINC 429 RCVR

NO.18

(H)

(L)

73

72

ARINC 429 RCVR

NO.17

(H)

(L)

89

75

99

100

ARINC 429 XMTR

NO. 4

(H)

(L)

RM-855 NO. 1(H)

(L)

AA

t

(H)

(L)

144J1

AAt

30

31

C144J1

FMS CSDBBUS IN

FMS CSDBBUS IN

RM-855 NO. 2

(H)

(L)TO GPWS

10111252

232425

3738

39

28 V DCPOWER

POWERGROUND

CHASSISGROUND

SIGNAL GROUND

AFIS

28V DC AM DC BUS 1B

(H)

(L)

CSDB FMSBUS XMTR

TO GNSSU

SIGNALGROUND


AD-49864-R1@

C9J1

149J1

Figure 2-8-5. FMS Interface Diagram



4. Fault Monitoring

A major failure within the FMS computer will result in numerous indications throughout thecockpit and possibly the CDU going dark. If, with the CDU dark, pressing several of thekeyboard buttons shows characters in the scratchpad area, the CDU is okay, and the FMScomputer is faulty. If pressing buttons does not cause characters to appear, then the CDU isfaulty.

In the event of minor failures, appropriate displays will be removed from the PFD and MFD.Note that some indications are removal of displays, while others cause the digital readouts to bechanged from numbers to dashes.

A. PFD Indications

Failures on the PFD are listed in Table 2-8-8. Figure 2-8-6 shows some of these failures.

Note that the PFD failure displays will not be observed if the FMS has not been selectedfor display via the DC-550 Display Controller.

Table 2-8-8. PFD Failure Indications

Invalid Indication

Vertical Deviation Removal of vertical deviation pointer and deviationscale

FMS Groundspeed Groundspeed digital readout = dashes

Lateral Deviation Data Removal of lateral deviation bar and lateral deviationscale to be red X’d

Desired Track Data Removal of desired track pointer and desired trackdigital readout = amber dashes

FMS Bearing Data Removal of FMS bearing pointer

FMS Distance to Waypoint Data Removal of waypoint identifier and distance towaypoint digital readout = amber dashes



220

LNAVAP

10 10

20 20

10 10

20 20

310 0031500

31020

8000

29.92 IN

DTK- - -

HDG050

0

1

23

32

1

N

6

15

24

S

W

E

GSPD- - - KTS

- - -. - NM

ALT

AD-51617-R1@

30500

.630 M

FMS

240

200

210

180

Figure 2-8-6. PFD FMS Failure Indications



B. MFD Indications

Failures on the MFD are listed in Table 2-8-9. Figure 2-8-7 shows some of these failures.

Table 2-8-9. MFD Failure Indications

Invalid Indication

Wind Information Removal of the wind display

FMS Distance to Waypoint Data Removal of waypoint identifier and distance towaypoint digital readout = amber dashes

FMS Time To Go TTG digital readout = dashes

FMS Drift Angle Data Removal of drift bug

Waypoint Data Removal of all waypoint symbols

NAVAID Data Removal of all NAVAID symbols

Airport Data Removal of all airport symbols

Holding Pattern Data Removal of the racetrack symbol

Flight Plan Data Removal of all flight plan informationVertical profile window and aircraft symbol remain

All FMS Data Removal of the flight plan designatorRemoval of the bearing/distance digital readoutRemoval of the LAT/LONG digital readoutRemoval of the navigation source indicator



N

S

33

15

3012

WE

246

21

3

300+15- - -

TASSATGSPD

FMS

- -. -- - -

NMMIN

360

50 50

TGTTX

SETV2- - -

VR110

V189RTN

SYSTEMDISPLAY

AREA

MAG1

AD-51618@

Figure 2-8-7. MFD FMS Failure Indications




SYSTEMMAINTENANCEMANUALFALCON 900EX

2.9 GLOBAL POSITIONING SYSTEM (GPS) (Optional) . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-1


A. Global Positioning System Sensor Unit . . . . . . . . . . . . . . . . . . . . . . 2-9-2

B. CD-810 Control Display Unit (CDU) GPS Status . . . . . . . . . . . . . . . . 2-9-3

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-6

A. Modes of Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-6

(1) Self-Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-6

(2) Initialization Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-6

(3) Acquisition Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-6

(4) Navigation (NAV) Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-7

(5) Aided Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-7

(6) Altitude Aiding Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-8

(7) Fault Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-8

(8) Mode Provisioning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-8

B. GPS Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-8

C. ARINC 429 Input Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-10

D. ARINC 429 Output Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-11





FIGURE/TITLE PAGE


Figure 2-9-1. Global Positioning System Sensor Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-2



Figure 2-9-4. Global Positioning System Interface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-9

List of Tables

TABLE/TITLE PAGE

Table 2-9-1. Global Positioning System Sensor Unit Leading Particulars . . . . . . . . . . . . 2-9-2

Table 2-9-2. GNSSU ARINC 429 Input Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-10

Table 2-9-3. GNSSU ARINC 429 Output Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9-11



SECTION 2.9GLOBAL POSITIONING SYSTEM (GPS) (Optional)

1. General

The PRIMUS® 1000 Integrated Avionics System installed in the Embraer 145 aircraft includes anoptional global positioning system (GPS). The GPS is used to determine aircraft position fromNAVSTAR satellites and provides this position data to the flight management system.

The GPS consists of the following LRUs:

• Global Navigation System Sensor Unit (GNSSU)

• Antenna (Non Honeywell).




A. Global Positioning System Sensor Unit

Figure 2-9-1 shows a graphical view of the Global Positioning System Sensor Unit. It islocated in the forward cabin avionics cabinet. Table 2-9-1 provides items andspecifications that are particular to the unit.

AD-40902@

Figure 2-9-1. Global Positioning System Sensor Unit

Table 2-9-1. Global Positioning System Sensor Unit Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . 2.50 in. (63.5 mm)

• Width . . . . . . . . . . . . . . . . . 9.50 in. (241.3 mm)

• Length . . . . . . . . . . . . . . . . 7.50 in. (190.5 mm)

Weight . . . . . . . . . . . . . . . . . . . . 5.92 lb (2.69 kg)


User Replaceable Parts . . . . . . . . None

Mating Connector . . . . . . . . . . . . M83723/77R2041N

Mounting . . . . . . . . . . . . . . . . . . Hard Mount Using Four 10-32 Cap Screws



The GNSSU is a flange-mounted LRU with two connectors. One connector is the interfaceto the aircraft wiring harness and the other is a coaxial connector for the antenna cable.

The GNSSU is a 12-channel GPS receiver that receives the L1 transmissions from theNAVSTAR satellite constellation. Using these signals, the GNSSU computes the antennaposition and then outputs data to the FMS. The output data includes three dimensionalaircraft position and velocities, satellite position, pseudo range, and delta range data.

The GNSSU does not have a controller. Its operation and mode selection is fully selfcontrolled. When the FMS is active, the CDU allows the operator access to several pagesof display data.

The GNSSU is provisioned to accommodate a number of functions which are evolving oranticipated as additions to the GNSSU functions. Most of these changes will beincorporated into the GNSSU through software changes only, and no hardwaremodifications will be required. Others may require changes to, or addition of GNSSUhardware.

The changes which Honeywell will supply through data loader installed software updatesare:

• Differential GPS

• RAIM (Receiver Autonomous Integrity Monitor) enhancements.

The GNSSU will require hardware upgrades for the following functions:

• Global Orbiting Navigation Satellite System (GLONASS) satellite signal reception

• GLONASS SATCOM signal rejection filtering.

Full details of the operation of the FMS CDU is given in SECTION 2.8.

B. CD-810 Control Display Unit (CDU) GPS Status

See Figures 2-9-2 and 2-9-3.

When requested by the operator, the CDU displays GPS STATUS. GPS STATUS pagesare accessed as follows:

• Press the NAV function button on the FMS CDU to activate the NAV INDEX displaypage

• Press the line select key adjacent to the POS SENSORS arrow to activate the POSSENSORS display page

• Press the line select key adjacent to the STATUS arrow on the GPS 1 line to activatethe GPS 1 STATUS 1/2 page (see Figure 2-9-2).

NOTE: Pressing either the NEXT or PREV function keys on the CDU will togglebetween pages GPS STATUS 1/2 and GPS STATUS 2/2 (see Figure 2-9-3.)



AD-43141@

Figure 2-9-2. GPS STATUS Page 1/2

AD-43142@

Figure 2-9-3. GPS STATUS Page 2/2



The GPS STATUS 1/2 page, shown in Figure 2-9-2, shows the following information:

• GPS position

• Groundspeed

• Altitude

• Miles from FMS position.

GPS altitude is the absolute altitude above the earth and should approximate BAROaltitude.

NOTE: GPS altitude and BARO altitude can differ by as much as 400 feet.

GPS STATUS 2/2 page, shown in Figure 2-9-3, displays the following information:

• Figure of Merit (FOM)

• Horizontal Dilution of Precision (HDOP)

• Vertical Dilution of Precision (VDOP)

• Time (UTC) and Date

• Operating Mode

• Satellites Tracked

• Satellites Viewed.

FOM indicates a position of uncertainty while HDOP and VDOP generate informationregarding satellite geometry. As a general rule, the smaller the number for FOM, HDOP,and VDOP, the better the accuracy of position.

The fourth line displays the operational modes of the GPS. The operational modes thatcan be displayed are:

• SELF-TEST

• INITIALIZATION

• ACQUISITION

• NAVIGATION

• ALTITUDE AIDING

• FAILED.



3. Operation

A. Modes of Operation

The GNSSU has seven operational modes:

• SELF-TEST

• INITIALIZATION

• ACQUISITION

• NAVIGATION (GPIRS sensors display AUTONOMOUS NAV or HYBRID rather thanNAVIGATION)

• AIDED

• ALTITUDE AIDING

• FAULT.

(1) Self-Test Mode

The GNSSU is in the Self-Test Mode for a maximum of 5 seconds from when itreceives power until it completes all internal power-up built-in tests (BITs). While itis in the Self-Test Mode, the GNSSU does not output data on the ARINC 429 BUS.When it completes the Self-Test Mode, the GNSSU enters either the InitializationMode or the Fault Mode.

(2) Initialization Mode

The GNSSU automatically enters the Initialization Mode to initialize its hardwarewhen it completes the Self-Test Mode. It is in this mode for only a fraction of asecond. When the hardware is initialized, the GNSSU enters the Acquisition Mode.

(3) Acquisition Mode

The GNSSU enters the Acquisition Mode from the Initialization Mode to acquiresatellites, or from other modes (NAV or Aided) when it does not have enoughsatellite and/or aiding data to remain in either the NAV or the Aided Mode. From theAcquisition Mode, the GNSSU enters either the NAV or Fault Mode. The AcquisitionMode proceeds in a number of ways.

(a) The GNSSU acquires satellites based on the information that it has when itenters the Acquisition Mode. To acquire satellites, the GNSSU uses this data:

• Almanac data that provides the coarse satellite orbits. The GNSSU storesalmanac data in non-volatile memory, which does not require an internal orexternal battery for operational support.

• Time , which is used with almanac data, to estimate the present position ofthe satellites and their orbits. The GNSSU may receive time and date fromthe FMS on the ARINC 429 input bus, or by acquiring a satellite.

• The approximate GNSSU location , which helps to predict which satellitesare visible. The GNSSU may receive position data from the FMS on theARINC 429 input bus.



(b) When the GNSSU has the information that is necessary to acquire satellites, itpredicts which satellites are visible and then acquires the satellite signals. Itcollects ephemeris data by decoding the satellite down-link data message.Ephemeris data is precise orbital data for a particular satellite. When itacquires each satellite, the GNSSU begins to transmit the satellitemeasurement data for that satellite. When it is tracking a sufficient number ofsatellites, the GNSSU computes position and velocity and enters the NAVMode.

(c) When the GNSSU does not have almanac and/or initialization data, it does aSearch the Skies acquisition. To do this, the GNSSU attempts to acquire all ofthe satellites in the GPS constellation. When it acquires the first satellite, itdecodes the satellite’s ephemeris data from a down-link message. When ithas acquired a sufficient number of satellites, the GNSSU enters thenavigation mode. Without valid initial data, the time-to-first-fix (TTFF) of asatellite is less than 10 minutes. With initialization and almanac dataavailable, the TTFF of a satellite is less than 75 seconds (95% confidencelevel).

(4) Navigation (NAV) Mode

The GNSSU enters the NAV Mode when it has computed a navigation solution thatprovides position, velocity, and time measurements. The GNSSU enters the NAVMode from the Acquisition Mode or from the Aided Mode. From the NAV Mode, theGNSSU enters the Acquisition, Aided, or Fault Mode.

(5) Aided Mode

The GNSSU enters the Aided Mode when insufficient satellite and/or altitudeinformation is available but external-aided data is available to continue to update theNAV filter. The GNSSU enters this mode from the NAV Mode, and while in thismode, provides valid time outputs. From this mode, the GNSSU enters the NAV,Acquisition, or Fault Mode.

(a) This mode may use inertial velocities to aid the navigation solution andintegrity monitoring during extended periods of insufficient satellite coverageand geometry.

(b) The GNSSU may enter the Aided Mode only when there are insufficientsatellites tracked to remain in the NAV Mode.

The GNSSU will remain in this mode for a maximum of 5 minutes.



(6) Altitude Aiding Mode

If satellite measurements are not sufficient for the GPS sensor to maintain integrityor remain in the NAV Mode, yet are sufficient when altitude information is available,the GNSSU is in Altitude Aiding mode. This mode uses altitude data from the MADCto aid the navigation solution and integrity monitoring during extended periods ofinsufficient satellite coverage and geometry. The GNSSU enters Altitude AidingMode only after the pressure altitude has been calibrated with a geometric altitudesolution using GPS with sufficient integrity. When the calibrated pressure altitudestandard deviation estimate is out of limits, it reverts to the Aided Mode. AltitudeAided Mode is entered from the NAV or Aided Modes, and exits to the NAV, Aided,or Fault Modes.

(7) Fault Mode

The GNSSU enters the Fault Mode when its outputs are affected by one or morecritical system faults. The GNSSU enters this mode from any other mode. Thismode supersedes all other modes or operation, and remains active until the nextpower-up cycle.

(8) Mode Provisioning

The GNSSU is provisioned for the Differential Mode. However, this mode is notcurrently implemented in equipment certified under TSO C129.

B. GPS Interface

See Figure 2-9-4.

The GNSSU accepts position data from the FMS via an ARINC 429 bus. This data is usedby the GNSSU during the acquisition mode, and is reported back on an ARINC 429 bus asFMS data. Once the GNSSU enters the navigation mode, the data is reported as GPSposition data and the FMS uses the data for its navigation purposes.

The DL-900 Data Loader can be used to install software updates.

WOW is used by the GNSSU to determine which data to process from the FMS, becausethe data is different with and without WOW.

The GNSSU also accepts altitude data from the MADCs, via ARINC 429, during theAltitude Aided Mode.

Other connections include power and ground.



HL

HL

AZ-850 MADC NO. 1

AZ-850 MADC NO. 2

NZ-2000 NAV COMPUTER (Optional)

1718

HL

HL

NO. 1 ARINC 429 IN

FMS GENERALBUS NO. 1

NO. 4 ARINC 429 OUT

NO. 4 ARINC 429 OUT

121J1

9J1

C9J1

34

HL

HL

NO.1 ARINC 429 OUT

NO. 1 ARINC 429 FMS IN

+28V DC GROUND

+28V DC POWER IN

ARINC 429NO. 1 MADC IN

ARINC 429NO. 2 MADC IN

35

67

HL

HL

+28V DCPOWER IN

149J1

38

39

18

37

10

11

68

69

68

69

62

49

A B

DL-900 DATALOADER (Optional)123J1

Z

a

E

RS-232 RCV DATA

RS-232 XMIT DATA

LOADER CONNECTEDLEFT

RS-232 XMTR

RS-232 RCVRDATA LOADER

MODE DISCRETE

2

3

12

W.O.W. 28

POWERGROUND

33CHASSIS GROUND CHASSISGROUND8

21ARINC 429 LO SPD

OUTPUT SELECT

SDI NO. 1 36 SIGNALGROUND

GLOBAL NAVIGATIONSYSTEM SENSORUNIT (Optional)

AD-49860-R1@

Figure 2-9-4. Global Positioning System Interface



C. ARINC 429 Input Data

The GNSSU accepts ARINC 429 low speed data in three different data formats as listed inthe following table. The GNSSU can accept and process all of the 429 labels listed inTable 2-9-2. Availability of these labels from the FMS depends on whether AHRS or IRS ison board. Air data parameters are from the MADC.

Table 2-9-2. GNSSU ARINC 429 Input Data

OCTALLABEL

PARAMETER/SIGNAL NAME UNITS DIGITALRANGE

RESOLUTION

BINARY (BNR) DATA FORMAT

131 Alternate Wpt Longitude Degrees ± 180° 0.000688

132 Alternate Wpt Latitude Degrees ± 180° 0.000688

135 Alternate Wpt E.T.A. HR:MN 23:59 N/A

143 Destination Wpt Longitude Degrees ± 180° 0.000688

144 Destination Wpt Latitude Degrees ± 180° 0.000688

150 UTC HR:MN:S 23:59:9 N/A

152 Destination Wpt E.T.A.

203 Altitude Feet ± 131,072 1

210 True Air Speed Knots 2048 0.0625

310 Latitude Degrees ± 180° 0.000172

311 Longitude Degrees ± 180° 0.000172

312 Ground Speed Knots 4096 0.125

313 Track Angle True Degrees ± 180° 0.0055

314 True Heading Degrees ± 180° 0.0055

324 Pitch Angle Degrees ± 180° 0.0055

325 Roll Angle Degrees ± 180° 0.0055

361 Altitude - Inertial Feet ± 131,072 0.125

365 Vertical Speed Feet/Min ± 32,768 1

BINARY CODED DECIMAL (BCD) DATA FORMAT

040 Set Altitude Feet 100,000 1

041 Set Latitude Deg/Min ± 180° 0.1 Min

042 Set Longitude Deg/Min ± 180° 0.1 Min

125 UTC HR:MIN 23:59.9 0.1 Min

260 Date D:M:YR N/A 1 Day

DISCRETE (DIS) DATA FORMAT

126 SV Deselect A N/A N/A N/A

127 SV Deselect B N/A N/A N/A



D. ARINC 429 Output Data

The GNSSU provides ARINC 429 high speed output data in three different data formats aslisted in Table 2-9-3. The output data is updated approximately once per second. Forfurther information regarding ARINC 429, refer to SECTION 1 of this manual. Specificdigital word format can be found in the Installation Manual for the GNSSU, HoneywellPublication No. M15-3819-009.

Table 2-9-3. GNSSU ARINC 429 Output Data

OCTALLABEL


RESOLUTION


061 Pseudo Range Meters ± 268435456 256

062 Pseudo Range Fine Meters 256 0.125

063 Pseudo Range Rate Meters/Sec ± 4096 0.0039

064 Delta Range Meters ± 4096 0.0039

065 Satellite Position X Meters ± 67108864 64

066 Satellite Position X Fine Meters 64 0.0039

070 Satellite Position Y Meters ± 67108864 64

071 Satellite Position Y Fine Meters 64 0.0039

072 Satellite Position Z Meters ± 67108864 64

073 Satellite Position Z Fine Meters 64 0.0039

074 UTC Measured Time Seconds 10.0 9.536743 µSec

076 GPS Altitude (MSL) Feet ± 131072 0.125

101 HDOP N/A 1024 0.031

102 VDOP N/A 1024 0.031

103 GPS Track Angle - True Degrees ± 180 0.0055

110 GPS Latitude Degrees ± 180 0.000172

111 GPS Longitude Degrees ± 180 0.000172

112 GPS Ground Speed Knots 4096 0.125

120 GPS Latitude Fine Degrees 0.000172 8.38E-8

121 GPS Longitude Fine Degrees 0.000172 8.38E-8

125 UTC HR:MIN 23:59:9 0.1 Min

130 Aut. Horiz Integrity Limit NM 16 1.2E-4

133 Aut. Vert Integrity Limit Feet 32768 0.125

135 Approach Area VIL Feet 32768 0.25

136 Vertical FOM Feet 32768 0.125

140 UTC Fine Seconds 1.0 .9536743 µSec

141 UTC Fine Fractions Seconds .953674 µSec .9313225 nSec

143 Approach Area HIL NM 16 0.00012

150 UTC HR:MM:SS 23:59:59 1 Sec



Table 2-9-3. GNSSU ARINC 429 Output Data

OCTALLABEL


RESOLUTION


162 Destination Waypoint E.T.A. HR:MM 23:59 1 Min

163 Alternate Waypoint E.T.A. HR:MM 23:59 1 Min

165 Vertical Velocity Feet/Min ± 32768 1.0

166 N/S Velocity Knots ± 4096 0.125

174 E/W Velocity Knots ± 4096 0.125

247 Horizontal FOM Feet 16 6.1E-5

343 Destination Waypoint HIL NM 16 0.0078

347 Alternate Waypoint HIL NM 16 0.0078

BINARY CODED DECIMAL (BCD) DATA FORMAT

260 Date D:M:YR N/A 1 Day

377 Equipment ID N/A N/A N/A

DISCRETE (DIS) DATA FORMAT

156 Maintenance

157 Maintenance

273 GPS Sensor Status N/A N/A N/A

352 Maintenance Discrete 1

354 System Time Counter Seconds 262144 1 Sec

355 Maintenance Discrete 2



4. Fault Monitoring

Any fault indications within the GPS will be reported on the FMS CDU. The MSG indicator onthe CDU will illuminate, and GPS FAILED will be displayed in the scratchpad area.



(Blank Page)





2.10 FLIGHT DIRECTOR SYSTEM (FDS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-1

A. Flight Director Data Management . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-2

B. Flight Director Couple Switching . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-2

C. Master/Slave Air Data Target Switching . . . . . . . . . . . . . . . . . . . . . 2-10-2

D. Flight Director Mode Synchronization . . . . . . . . . . . . . . . . . . . . . . 2-10-3

E. Flight Director Mode Annunciation . . . . . . . . . . . . . . . . . . . . . . . . 2-10-3

F. Flight Director Command Bar Logic . . . . . . . . . . . . . . . . . . . . . . . 2-10-3

G. Altitude Preselect Function . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-3

2. Component Descriptions and Locations . . . . . . . . . . . . . . . . . . . . 2-10-4


B. GC-550 Guidance Panel Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-6

(1) Flight Director Mode Switches . . . . . . . . . . . . . . . . . . . . . . 2-10-8

(2) Heading (HDG) Select Knob . . . . . . . . . . . . . . . . . . . . . . . . 2-10-8

(3) Course (CRS) Select Knob . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-9

(4) Altitude Select Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-9

(5) Couple (CPL) Pushbutton . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-9

(6) Bank Pushbutton (BNK) . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-9

(7) FD1/FD2 Pushbuttons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-9

C. DC-550 Display Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-10

D. PC-400 Autopilot Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-12

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-14

A. Flight Director Functions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-14

(1) PFD Command Bars . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-14

(2) GS CAP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-15

(3) GS Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-15

(4) Lateral Beam Sensor (LBS) . . . . . . . . . . . . . . . . . . . . . . . 2-10-15

(5) LOC/BC CAP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-15

(6) LOC/BC TRACK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-15

(7) True Airspeed (TAS) Gain Programming . . . . . . . . . . . . . . 2-10-16

(8) Vertical Beam Sensor (VBS) . . . . . . . . . . . . . . . . . . . . . . . 2-10-16

(9) Vertical Path Gain Programming . . . . . . . . . . . . . . . . . . . 2-10-16

(10) VOR CAP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-16

(11) VOR Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-16

(12) VOR OSS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-17

(13) VOR AOSS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-17





B. Flight Director Lateral (Roll) Channel Functional Operation . . . . . 2-10-17

(1) Flight Director Lateral (Roll) Modes Interface . . . . . . . . . . 2-10-17

(2) Heading Select (HDG) Mode . . . . . . . . . . . . . . . . . . . . . . . 2-10-27

(3) Heading Select Mode Engage/Reset/Disengage Logic . . . . 2-10-28

(4) VOR (NAV) Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-29

(5) VOR Approach (VAPP) Mode . . . . . . . . . . . . . . . . . . . . . . 2-10-36

(6) VOR/VAPP Engage/Reset/Disengage Logic . . . . . . . . . . . . 2-10-37

(7) Localizer (NAV) and Back Course (BC) Modes . . . . . . . . . 2-10-41

(8) Localizer/Back Course Mode Engage/Reset/Disengage

Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-51

(9) Long Range Navigation . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-54

(10) LNAV Mode Engage/Reset/Disengage Logic . . . . . . . . . . . 2-10-57

C. Flight Director Vertical (Pitch) Channel Functional Operation . . . 2-10-59

(1) Flight Director Vertical (Pitch) Modes Interface . . . . . . . . 2-10-59


(3) Pitch Attitude Hold Mode Engage/Reset/Disengage Logic . 2-10-70

(4) Vertical Speed (VS) Hold Mode . . . . . . . . . . . . . . . . . . . . . 2-10-71

(5) Vertical Speed (VS) Hold Mode Engage/Reset/Disengage

Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-73

(6) Speed (SPD) Select Mode . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-74

(7) Speed (SPD) Select Mode Engage/Reset/Disengage Logic 2-10-76

(8) Flight Level Change (FLC, FLCH) Mode . . . . . . . . . . . . . . 2-10-77

(9) Flight Level Change (FLC) Mode Engage/Reset/Disengage

Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-79

(10) Altitude Preselect (ASEL) Mode . . . . . . . . . . . . . . . . . . . . 2-10-80

(11) Altitude Preselect (ASEL) Mode Engage/Reset/Disengage

Logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-87

(12) Altitude Hold (ALT) Mode . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-88

(13) Altitude Hold (ALT) Mode Engage/Reset/Disengage Logic 2-10-90

(14) ILS Approach (APR) Mode . . . . . . . . . . . . . . . . . . . . . . . . 2-10-91

(15) ILS Approach (APR) Mode Engage/Reset/Disengage Logic 2-10-97

(16) Go-Around (GA) Mode (Wings Level) . . . . . . . . . . . . . . . . 2-10-99

(17) Go-Around (GA) Mode Engage/Reset/Disengage Logic . . . 2-10-101

(18) Windshear Mode (WSHR) . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-102

(19) Windshear (WSHR) Mode Engage/Reset/Disengage Logic . 2-10-103




FIGURE/TITLE PAGE



Figure 2-10-2. GC-550 Guidance Control Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-6

Figure 2-10-3. DC-550 Display Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-10


Figure 2-10-5. Flight Director Lateral Modes Interface - Pilot’s Side . . . . . . . . . . . . 2-10-23/24

Figure 2-10-6. Flight Director Lateral Modes Interface - Copilot’s Side . . . . . . . . . . 2-10-25/26

Figure 2-10-7. VOR ARM Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-29

Figure 2-10-8. VOR (NAV) Mode Armed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-30

Figure 2-10-9. VOR Capture Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-31

Figure 2-10-10. VOR (NAV) Mode Capture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-32

Figure 2-10-11. VOR Course Cut Limiting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-33

Figure 2-10-12. VOR Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-34

Figure 2-10-13. VOR Overstation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-35

Figure 2-10-14. Localizer ARM Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-41

Figure 2-10-15. Localizer (NAV) Mode ARM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-42

Figure 2-10-16. Localizer Capture Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-43

Figure 2-10-17. Localizer Capture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-45

Figure 2-10-18. Localizer Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-46

Figure 2-10-19. Back Course Mode Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-47

Figure 2-10-20. Back Course Arm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-48

Figure 2-10-21. Back Course Capture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-49

Figure 2-10-22. Back Course Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-50

Figure 2-10-23. Long Range Navigation Capture Pictorial and Tracking . . . . . . . . . . . . 2-10-55

Figure 2-10-24. Long Range Navigation Tracking . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-56

Figure 2-10-25. Flight Director Vertical Modes Interface - Pilot’s Side . . . . . . . . . . . 2-10-65/66

Figure 2-10-26. Flight Director Vertical Modes Interface - Copilot’s Side . . . . . . . . . 2-10-67/68

Figure 2-10-27. Vertical Speed (VS) Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-72

Figure 2-10-28. Speed Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-75

Figure 2-10-29. FLC Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-78

Figure 2-10-30. Altitude Preselect Mode Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-82

Figure 2-10-31. Prior to Descent - Altitude Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-83

Figure 2-10-32. During Descent - ASEL Armed Mode . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-84

Figure 2-10-33. Start of Flare - ASEL Capture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-85

Figure 2-10-34. Level at New Altitude - Altitude Hold Mode . . . . . . . . . . . . . . . . . . . . . 2-10-86

Figure 2-10-35. Altitude Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-89

Figure 2-10-36. ILS Approach Arm Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-92




FIGURE/TITLE PAGE


Figure 2-10-37. ILS Approach Arm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-93

Figure 2-10-38. ILS Approach Capture Pictorial . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-94

Figure 2-10-39. ILS Approach (LOC) Capture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-95

Figure 2-10-40. ILS Approach (APR) Mode Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-96

Figure 2-10-41. Go-Around Mode (Wings Level) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-100

Figure 2-10-42. Windshear Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-103

List of Tables

TABLE/TITLE PAGE

Table 2-10-1. IC-600 Integrated Avionics Computer Leading Particular s . . . . . . . . . . . . 2-10-5

Table 2-10-2. GC-550 Guidance Panel Unit Leading Particulars . . . . . . . . . . . . . . . . . . . 2-10-6

Table 2-10-3. DC-550 Display Controller Leading Particulars . . . . . . . . . . . . . . . . . . . . 2-10-10


Table 2-10-5. Heading Select Mode Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-27

Table 2-10-6. VOR/VOR Approach Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-36

Table 2-10-7. Localizer (LOC) and Back Course (BC) Mode Operating Limits . . . . . . . . 2-10-51

Table 2-10-8. LNAV Mode Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-57

Table 2-10-9. Pitch Attitude Hold Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-69

Table 2-10-10. Vertical Speed Hold Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . 2-10-72

Table 2-10-11. Speed (SPD) Hold Mode Operating Limits . . . . . . . . . . . . . . . . . . . . . . . 2-10-76

Table 2-10-12. Flight Level Change (FLC) Hold Mode Operating Limits . . . . . . . . . . . . 2-10-79

Table 2-10-13. Altitude Preselect (ASEL) Mode Operating Limits . . . . . . . . . . . . . . . . . 2-10-87

Table 2-10-14. Altitude Hold (ALT) Mode Operating Limits . . . . . . . . . . . . . . . . . . . . . 2-10-89

Table 2-10-15. ILS Approach (APR) Mode Operating Limits . . . . . . . . . . . . . . . . . . . . . 2-10-96



SECTION 2.10FLIGHT DIRECTOR SYSTEM (FDS)

1. General

The PRIMUS® 1000 Display and Flight Guidance System employs two separate flight directorsystems. One is housed in the pilot’s IC-600 Integrated Avionics Computer (IAC) and the otheris housed in the copilot’s IC-600 IAC. The flight directors provide computed steering commandsto the autopilot and to the command bars on the PFDs. With the autopilot not engaged, thepilot can manually fly the steering command. With the autopilot engaged, the flight directorcomputed steering commands will be flown by the autopilot. The flight director provides bothlateral and vertical steering commands and one each can be active at the same time to controlthe aircraft. Other flight director modes can be armed to automatically become active at theproper time.

Each flight director system consists of the following components:

• GC-550 Guidance Control Unit (common to both flight directors)

• IC-600 Integrated Avionics Computer (IAC)

• PC-400 Autopilot Controller

• DC-550 Display Controller.

The sensors utilized by the flight director are:

• AZ-850 Micro Air Data Computer

• AH-800 Attitude Heading Reference Unit

• RNZ-851 Integrated Navigation Unit

• Long Range Navigation System

• AA-300 Radio Altimeter.

For the flight director to compute a steering command, the following has to be considered:

• What is the pilot’s desired attitude/position?

• What is the aircraft’s actual attitude/position?

• If there is a difference between desired and actual, correct for the difference and control thespeed at which the correction takes place.

Flight director modes utilize on-side attitude data for computations. Cross-side attitude data isused for monitoring only.



A. Flight Director Data Management

The flight director only requires pitch and roll attitude for synchronizing the attitudecommand and for computing command bar outputs. The on-side attitude data is used forsynchronization and command bar computations when modes are selected. The IC-600IAC uses the flight director attitude command to position the command bar on the attitudesphere.

The flight director can couple to either short range NAV (SRN) or the long range NAV(FMS), based on which is being displayed on the PFD. Each flight director uses on-sidedisplayed NAV data.

The flight director uses the air data input as the reference for all vertical modes exceptglideslope, and for gain programming. The altitude hold target, airspeed target, verticalspeed target and selected altitude are all computed by the IC-600 IAC.

B. Flight Director Couple Switching

Flight director couple switching allows either the pilot’s or copilot’s flight director to becoupled to the PRIMUS® 1000 autopilot. This is accomplished by means of the transfer(CPL) pushbutton on the GC-550 Guidance Control Panel. The pilot’s IC-600 IAC controlsthe state of the CPL pushbutton discretes from the GC-550 Guidance Control Unit andcontrols the state of the flight director couple switching by means of the IC Bus. The pilot’sIC-600 IAC outputs the annunciator drive to the left arrow annunciator on the GC-550. Thecopilot’s IC-600 IAC outputs the annunciator drive to the right arrow annunciator on theGC-550. The power up state is left.

Activation of the CPL pushbutton resets the selected flight director modes of both flightdirectors. The pilot must then re-engage the flight director modes he wants active.

Activation of the on-side SG reversionary function automatically forces the CPL couple tothe cross-side flight director, if the on-side flight director is the master. Forced flightdirector switching will automatically reset the selected modes of both flight directors andactivate the left or right arrow annunciator depending on which flight director becomes themaster. If the on-side SG reversion function is activated on the side with a slaved flightdirector, the flight director modes on both flight directors will remain active.

C. Master/Slave Air Data Target Switching

The flight directors operate in a master/slave arrangement to allow tracking of the air datatargets, the heading select reference and the selected mode annunciations on the PFDbetween the pilot’s and copilot’s systems. The status of the GC-550 CPL couple indicateswhich flight director is the master. The master flight director computes the reference forthe selected altitude, SPD hold target, and vertical speed mode reference for both flightdirectors. This capability prevents vertical command bar splits between the master andslaved flight directors. When the master and slaved flight directors are in the same mode,the slaved flight director will synchronize to the targets computed by the master flightdirector.



D. Flight Director Mode Synchronization

The flight director modes are synchronized in a manner that allows control of the mode tobe accomplished by the single set of mode select buttons on the GC-550 Guidance ControlUnit. The flight director couple status determines which flight director is master for modeengagement. The slaved flight director will attempt to engage the same mode(s) as longas the appropriate valids and NAV sources are available. If the mode cannot be engagedon the slaved flight director, the slaved command bars will drop until such time as themode can be entered automatically.

E. Flight Director Mode Annunciation

The master flight director will annunciate the selected mode on both PFDs and thecross-side flight director will slave to this display.

If the IC Bus becomes invalid, each flight director will default to the master state and theflight director mode annunciations will be independently annunciated on the on-side PFD.When both flight directors revert to the master state, the GC-550 will default to the leftflight director.

After an IC Bus failure, it will be impossible to couple the right flight director to theautopilot. The right flight director will still be capable of accepting mode select inputs fromthe GC-550 and displaying annunciations on the PFD.

F. Flight Director Command Bar Logic

The flight director command bar (either cross pointer or single cue) are biased out of viewfor invalid data sources related to the mode. The flight director command bars may bemanually biased out of view by toggling the FD1 or FD2 buttons on the GC-550 GuidanceControl Unit.

G. Altitude Preselect Function

The altitude preselect function consists of a digital set knob located on the GC-550Guidance Control Unit. There also are altitude preselect light and horn warning discretes,controlled by the IC-600 IAC. The manual preselect function allows the pilot to select adesired altitude reference. The altitude set knob interfaces with the DC-550 DisplayController. Rotating the altitude set knob results in 100 foot increments of altitude targetvalue.

After power is applied to the system, the altitude display window on the PFD will indicatethree dashes in the middle digits, prior to turning the altitude set knob. When the knob isfirst turned, the altitude display window will indicate the present barometric altitudedisplayed on the pilot’s altimeter. The preselected altitude can be slewed up or down fromthe present barometric altitude value. A failed condition is indicated by amber dashes inthe display window.





Figure 2-10-1 shows a graphical view of the IC-600 Integrated Avionics Computer. TwoIC-600 Integrated Computers are located under the cockpit floor behind the crew.Table 2-10-1 provides items and specifications that are particular to the computer.

AD-33449@


The primary component of the flight director system is the IC-600 Integrated AvionicsComputer (IAC). Each IC-600 IAC is a symbol generator, fault warning computer, flightdirector and autopilot computer integrated into a single unit. All aircraft sensors andnavigation sources are connected directly to the IC-600 IAC, since all flight controlfunctions reside inside this computer.

Basic flight director modes are initiated by manual selection through the GC-550 GuidanceControl Unit. Once a mode is initiated, automatic transitions can occur from armed toactive status or to another mode if the transition initiation requirements are met. Thearmed mode states only provide a visual indication (PFD annunciation) of mode statusrelative to a manual selection of some guidance modes, whereas active mode statesprovide both visual mode status indications and pitch/roll steering commands to the PFDand the autopilot when engaged.



Data used to compute guidance commands are consistent with that displayed on the PFD.This data includes:

• Displayed heading and heading flag valid

• Selected course and course error

• Selected heading and heading error

• Lateral and vertical path deviations and flag valids

• DME distance, tuned-to-NAV and to-from status

• Middle marker data

• NAV source identification (tuned-to-localizer, VOR, LNAV)

• Lateral steering commands and flag valids.

Table 2-10-1. IC-600 Integrated Avionics Computer Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . 7.62 in. (193.55 mm)

• Width . . . . . . . . . . . . . . . 4.13 in. (104.90 mm)

• Length . . . . . . . . . . . . . . 16.45 in. (418.83 mm)

Weight (maximum):

• With Autopilot . . . . . . . . . 15.5 lb (7.05 kg)

• Without Autopilot . . . . . . . 15.0 lb (6.82 kg)

Power Requirements (with autopilot):



• Servo Power . . . . . . . . . . 28 V dc, 210 W (max)/112 W (nom)

Power Requirements (without autopilot):



Mating Connectors:

• J1, J2 . . . . . . . . . . . . . . . ITT Cannon Part No. DPX2MA-A106P-A106P-33B-0001Note: Sunbank backshell (4) required

Part No. J1560-12-2

Mounting . . . . . . . . . . . . . . . . Tray, HPN 7017095-903



B. GC-550 Guidance Panel Unit

Figure 2-10-2 shows a graphical view of the GC-550 Guidance Control Unit. It is located inthe center of the glareshield. Table 2-10-2 provides items and specifications that areparticular to the control panel.

CRS 1

PUSH SYNC PUSH SYNC

HDG SPD

PUSH IAS/M

ASEL CRS 2

PUSH SYNC

FD1 HDG

BNK

APR

NAV AP

CPL

YD

SPD FD2ALT

VS

FLC

AD-50630@

Figure 2-10-2. GC-550 Guidance Control Panel

Table 2-10-2. GC-550 Guidance Panel Unit Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.25 in. (57.15 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.65 in. (295.91 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.51 in. (114.51 mm)


Power Requirements:




Table 2-10-2. GC-550 Guidance Panel Unit Leading Particulars

Item Specification


• Knobs










• Setscrews



• Lamps



Mating Connectors:

• J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS27473E20B-35S




The GC-550 Guidance Control Unit provides flight director mode selection andde-selection, heading and course select as well as altitude set for the altitude pre-selectmode and speed set for the SPD and VS mode. The GC-550 also provides switchingcapability for selecting either the pilots or copilots flight director as the master. Flightdirector modes are:

Lateral Modes:

• HDG - Heading select

• NAV - Lateral navigation (SRN or LRN)

• APR - Localizer and Glideslope (also backcourse localizer)

• VAPR - VOR Approach

• GA - Go-Around (wings level).

• BNK - Bank (Heading Select only)

Vertical Modes:

• VS - Vertical Speed hold

• ALT - Altitude hold

• SPD - Indicated Airspeed/Mach hold

• FLC - Flight Level Change

• GS - Vertical guidance for ILS approach

• GA - Go-Around.

Automatic Mode:

• ASEL - Altitude Preselect (not selectable on the GC-550).

(1) Flight Director Mode Switches

Flight director mode selection is accomplished by seven mode select switches.Annunciation of active modes is provided by individual annunciators associated witheach mode select switch. These annunciations are illuminated for the armed andcaptured modes which correspond to each switch. The actual flight director modeannunciation is displayed on the PFD and this annunciation differentiates betweenarmed and captured modes.

(2) Heading (HDG) Select Knob

The single heading select knob sets the blue heading select bug on both PFDs. Thebug is positioned around the compass card in one degree increments. Fast knobrotation results in increased heading select bug motion. The heading select bug isused to set the desired heading when flying the flight director heading select mode.

The PUSH SYNC function allows the pilot to slew the heading bug to the compasscard fore lubber line.



(3) Course (CRS) Select Knob

This function is activated by two separate knobs on the GC-550 Guidance ControlUnit. The CRS 1 knob controls the pilot’s course select function and the CRS 2 knobcontrols the copilot’s course select function. Knob rotation allows course selection inone degree increments. Fast knob rotation results in increased course pointermotion. The course knob is used to set the selected course for VOR and localizerflight director modes.

The PUSH SYNC function allows the pilot to slew the course pointer on top of theappropriate bearing pointer to show the zero deviation path to the station.

(4) Altitude Select Knob

The altitude select knob is used to set a preselected altitude reference in feet in thealtitude display window on the PFD. This altitude reference is used by the flightdirector altitude preselect mode. Altitude can be set in hundreds and thousands offeet only.

(5) Couple (CPL) Pushbutton

This pushbutton is used to couple either the pilot’s or copilot’s flight director to thePRIMUS® 1000 autopilot. Only one flight director at a time can be coupled to theautopilot. Activation of the CPL pushbutton will cancel all flight director modesindependent of autopilot operation.

(6) Bank Pushbutton (BNK)

The bank pushbutton allows pilot selection of a reduced maximum bank angle for theheading select mode only. When active, the bank angle limits are reduced from ±27°to ±14°. The low bank angle limit is automatically selected when climbing through25,000 feet. Automatic canceling of low bank occurs when descending through24,750 feet.

(7) FD1/FD2 Pushbuttons

The primary function of these pushbuttons is to bias the FD command bars on or offof the PFDs. There are exceptions to this rule based on the engage status of theautopilot.

With the autopilot engaged, the coupled side FD command bars will always bedisplayed. The uncoupled side FD command bars will continue to be toggled on/offwith the appropriate FD pushbutton.

With the autopilot disengaged, pressing the FD1 or FD2 button only (no flightdirector modes are active), will not bring the FD command bars into view. Anysubsequent flight director mode selection will cause the FD command bars to bedisplayed. With the FD command bars in view on both PFDs, pressing the FD1 orFD2 pushbutton will remove the FD command bars from that side PFD only. Whenthe FD command bars are in view on one side only, pressing that side FD button willdisengage all selected FD modes.



C. DC-550 Display Controller

Figure 2-10-3 shows a graphical view of the DC-550 Display Controller. Two DC-550Display Controllers are located in the glareshield. Table 2-10-3 provides items andspecifications that are particular to the controller.

BRG BRG

NAV FMSET

OFF

NAV 1

ADF

GSPDTTG

FULLWX

FMS OFF

NAV 2

ADF

FMS

RA TEST

AD-50629-R1@

Figure 2-10-3. DC-550 Display Controller


Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.25 in. (57.15 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.75 in. (146.05 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.87 in. (174.50 mm)


Power Requirements:

• Primary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 V dc, 5.0 W (max)

• Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5 V ac, 5.0 W (max)




Item Specification


• Knobs

- BRG (Setscrew A) . . . . . . . . . . . . . . . . . . . HPN 7009437

- BRG ◊ (Setscrew A) . . . . . . . . . . . . . . . . . . . HPN 7009437

- RA (Setscrew B) . . . . . . . . . . . . . . . . . . . . . . HPN 7018748-1

- Test Switch HUB (Setscrew B) . . . . . . . . . . . . HPN 7009644-3

• Setscrews

- A (Multi-Spline, 2-56 x 1/8-inch, cup point) . . HPN 2500148-64

- B (Multi-Spline, 4-40 x 3/16-inch, cup point) . HPN 2500148-130

Mating Connectors:

• J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS27473E20-B35SB

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Standard Dzus Rail

The DC-550 provides a data acquisition function for the GC-550 Guidance Control Unit.Course and heading select inputs, as well as altitude preselect and SPD set data aretransmitted to the IC-600 IAC on a two wire digital bus. Leading particulars for the DC-550Display Controller are provided in Table 2-10-3.



D. PC-400 Autopilot Controller

The PC-400 Autopilot Controller provides the autopilot with TURN knob and PITCH wheelinputs. The controller is located on the pedestal with Dzus rails with four, one-quarter turnfasteners. Figure 2-10-4 shows a graphical view of the PC-400 Autopilot Controller.Leading particulars for the PC-400 Autopilot Controller are listed in Table 2-10-4.

TURNDESCEND

CLIMB

PITCH

AD-50485@

Figure 2-10-4. PC-400 Autopilot Controller



Table 2-10-4. PC-400 Autopilot Controller Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.625 in. (6.67 cm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.750 in. (14.60 cm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.150 in. (15.62 cm)


Power Requirements:

• Instrument Lighting . . . . . . . . . . . . . . . . . . . . . . . . . 0 to 5 V ac or dc

• Mode Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . +28 V dc


• Knob, Turn . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HPN 337136-1

• Setscrew, Bottom (Hex Socket, 8-32 x 5/8", cuppoint) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

HPN 0455-284

• Setscrew, Side (Hex Socket, 8-32 x 3/16", cup point) HPN 0455-274

• Lamp, Clear (Type 7341) . . . . . . . . . . . . . . . . . . . . HPN 0635-22

Mating Connector:

• J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS3116F20-41S

The pilot may input roll commands with the TURN knob or change pitch attitude referencewith the PITCH wheel with the autopilot engaged.

The TURN knob drives a center tap potentiometer and cam mechanism. The cammechanism provides a no command mechanical detent and command sensing (out ofdetent) signal to the IC-600 IAC. The PITCH wheel drives a tachometer generator whichprovides a CLIMB or DESCEND dc voltage output proportional to the rotation of the wheel.



3. Operation

A. Flight Director Functions

This paragraph provides an explanation for the IC-600 IAC major flight director functions,referred to throughout the remainder of this section. These functions are listedalphabetically and are titled as follows:

• PFD Command Bar

• GS CAP

• GS Track

• Lateral Beam Sensor (LBS)

• LOC/BC CAP

• LOC/BC Track

• True Airspeed (TAS) Gain Programming

• Vertical Beam Sensor (VBS)

• Vertical Path Gain Programming

• VOR CAP

• VOR Track

• VOR OSS

• VOR AOSS.

(1) PFD Command Bars

When a flight director steering command is applied to the command bar input, thebar (either crosspointer or single cue) will move left or right (roll), or up and down(pitch). This provides the required visual command for the pilot to maneuver theaircraft in the proper direction to reach and maintain the desired flight path.

If the information required to fly a lateral or vertical flight director mode becomesinvalid, the mode will either be canceled, or the command cue will be biased out ofview, dependent on what went invalid.



(2) GS CAP

The following conditions are necessary for glideslope capture:

• The localizer mode is either capture or track

• The vertical beam sensor (VBS) has tripped

• GS deviation is less than 0.5 degrees.

(3) GS Track

Glideslope track occurs after the flight director has captured the vertical path and isnow tracking the beam. The track phase provides for tighter flying of the glideslopebeam. The track mode occurs after the GS is captured plus 15 seconds.

(4) Lateral Beam Sensor (LBS)

When flying to intercept the VOR or localizer beam, the LBS determines the propertime for the mode to change from the arm to the capture phase of operation. TheLBS looks at course error, radio deviation, TAS and DME if available.

In simple terms, the LBS compares the magnitude of the course error to themagnitude of the radio deviation and takes TAS and DME into account in itscomputation. If radio deviation is larger in magnitude than course error, the mode isarmed. As the aircraft approaches the beam center, course error is constant andradio deviation is getting smaller. At some point, the magnitude of the radiodeviation is less than the magnitude of the course error. It is at this point that theLBS trips and the aircraft turns to line up on the VOR or localizer beam center.

If the intercept angle to the beam is very shallow, the LBS will not trip until theaircraft is near the beam center. For this reason, an override of the LBS occurswhen the beam deviation reaches a specified minimum to avoid beam standoff.

(5) LOC/BC CAP

Localizer (LOC) and Back Course (BC) capture will occur when the followingconditions are met:

• Lateral Beam Sensor has tripped

• Beam deviation is less than 0.75 degrees.

(6) LOC/BC TRACK

Localizer and back course track signify the aircraft being on beam center andcrosswind washout correction can take place. The track phase will occur when thefollowing conditions are met:

• LOC or BC is captured plus 4 seconds

• Localizer beam deviation is less than 1 degree



(7) True Airspeed (TAS) Gain Programming

TAS gain programming is used on the heading select, course select, pitch wheelcommand, air data commands (except IAS) and glideslope deviation. TAS gainprogramming is used to achieve the same aircraft response to flight guidancecommands, regardless as to the altitude and speed of the aircraft. The TAScomputation is derived from altitude, airspeed and outside air temperature.

(8) Vertical Beam Sensor (VBS)

When flying an ILS approach and the localizer is captured, the VBS determines theproper time for the glideslope mode to transition from arm to capture. The VBS isarmed as a function of the navigation radio being tuned to an ILS frequency andlocalizer is captured or track. The VBS looks at vertical speed, TAS and glideslopedeviation. The VBS trips when glideslope deviation is less than 150 mV and thecapture sensor is satisfied. The capture sensor looks at airspeed, glideslope beamrate of change and normal acceleration to determine the optimum capture point.

In the event the aircraft is paralleling the beam (no beam closure rate), the VBS willtrip at a vertical deviation less than 20 mV. This will reset the previously selectedpitch mode and change aircraft attitude to smoothly capture the glideslope beam.

(9) Vertical Path Gain Programming

Vertical path (glideslope) gain programming starts after the VBS trips. The gain isprogrammed as a function of radio altitude and vertical speed. If the radio altimeteris invalid, gain programming starts at GS capture and is controlled by a runwayheight estimator. The value estimated assumes a 1500 foot start point and gain ischanged as a function of TAS and time. At the middle marker the gain is set to apreset amount for the remainder of the approach.

(10) VOR CAP

VOR capture occurs when the LBS has tripped.

(11) VOR Track

VOR track will occur as the aircraft is established on beam center and the followingconditions are met:

• Lateral deviation rate is less than 50 feet/second

• Aircraft bank angle is less than 6 degrees.

At this time, crosswind correction is allowed to start if required.



(12) VOR OSS

The over station sensor (OSS) is used to detect the erratic radio signals encounteredin the area above the VOR ground station antenna. When these radio signals reacha predetermined level, they are no longer useful and OSS eliminates them from thecontrol law.

The VOR OSS trips when either one of the following conditions occur:

• Distance to the station is less than [0.25 barometric altitude/cosine 30 degrees]and DME is valid and not hold

• Lateral deviation is greater than 75 mV and beam rate is greater than8mV/second and DME not valid or DME is hold

(13) VOR AOSS

When the aircraft is flying in the OSS state, VOR beam deviation is constantlymonitored to determine when it again is useful and can be included in the controllaw. The after over station sensor (AOSS) does this monitoring. AOSS will occurwhen the following conditions are met:

• Beam deviation is less than 75 mV plus 20 seconds

• Beam rate is less than 25 feet/second.

NOTE: In the VOR Approach mode, beam deviation is less than 75 mVplus 4 seconds.

B. Flight Director Lateral (Roll) Channel Functional Operation

(1) Flight Director Lateral (Roll) Modes Interface

Figure 2-10-5 shows the LRU interface for the pilots side flight director lateralmodes. Figure 2-10-6 shows the LRU interface for the copilots side flight directorlateral modes.

The function of each LRU for each lateral mode is discussed in the followingparagraphs:

(a) AH-800 Attitude Heading Reference Unit (AHRU)

For all flight director lateral modes, the on-side AHRU provides actual aircraftroll attitude, rate of change, and magnetic heading references for the on-sideflight director. In the VOR and VOR Approach and Localizer modes, theon-side AHRU also provides a lateral acceleration term.

The cross-side AHRU provides the same terms to the IC-600 IAC, but thesesignals are used for EDS/flight director and autopilot monitoring purposes only.



(b) AZ-850 Micro Air Data Computer (MADC)

The on-side AZ-850 MADC provides the on-side IC-600 IAC with an ARINC429 input of air data values including true airspeed (TAS). The TAS signal isused in all lateral flight director modes for gain programming. The response ofthe aircraft should feel the same regardless of the aircraft’s airspeed andaltitude. Since it requires less flight control surface deflection at high speedand high altitude to complete a maneuver then it does at low speed and lowaltitude, changing the size of the signal as a function of TAS achieves thedesired results.

Should the AZ-850 MADC become invalid, a fixed bias TAS of 120 knots isused in the IC-600 IAC. The default value of TAS is set for the approachspeed region of flight.

(c) GC-550 Guidance Control Unit

The GC-550 Guidance Control Unit provides the means for the pilot toengage/disengage all lateral flight director modes. The GC-550 also providesthe controls to set the heading select bug for the heading select mode, as wellas setting the selected course in the VOR, VOR Approach, and Localizermodes.

The GC-550 provides button input data directly to the IC-600 IAC. Headingbug set and selected course data is transmitted to both DC-550 DisplayControllers. The DC-550 Display Controllers transmit the heading bug set andselected course data to both IC-600 IACs.

(d) DC-550 Display Controller

The DC-550 Display Controller provides an RS-422 digital bus interface(DC/SG Bus) between itself and the IC-600 IAC. Heading bug set andselected course inputs are routed through the DC-550 to be put on the digitalbus interface to the IC-600 IAC.

(e) RNZ-851 Integrated Navigation Unit

The RNZ-851 integrated NAV unit provides an RSB output of VOR, Localizerdeviation data, as well as marker beacon data. The DME receiver provides anRSB output as well as a DME enable and DME Hold discrete. The DMEsignal is used in the VOR and VOR Approach modes to gain program the VORsignal as a function of the aircraft approaching, or departing the VOR station.

(f) RT-300 Radio Altimeter

The radio altimeter provides an analog output of absolute altitude above theterrain. This signal is used by the flight director to gain program the localizersignal. Gain programming is required due to the directional qualities andbeam convergence characteristics of the localizer antenna.



As the aircraft approaches the runway, the localizer signal appears to getstronger and the beam appears to get narrower. By reducing the gain on thesignal as a function of the change in radio altitude, the computed steeringcommand will not take the aircraft out of the localizer beam envelope andreduces "S" turning.

Should the radio altimeter be invalid, localizer gain programming will start as afunction of glideslope capture and run down as a function of TAS and time. Atthe middle marker gain programming will be synchronized to a preset value.

(g) LNAV Unit

The long range navigation unit provides an ARINC 429 composite steeringcommand to the IC-600 IAC. This command is already gain programmed inthe LNAV unit and represents the computed desired track over the groundfrom the last sequenced waypoint to the "TO" waypoint.

(h) IC-600 Integrated Avionics Computer (IAC)

The IC-600 IAC performs the following as a function of what lateral mode isactive.

1 Heading Select Mode

When the heading select mode is activated, the flight director processor inthe IC-600 IAC compares actual aircraft heading against desired aircraftheading, as determined by the position of the heading select bug on thePFD. The difference is the heading select error signal.

With the autopilot not engaged, the heading select error signal ispresented on the PFD flight director command bar as a steering commandfor the pilot to bank the aircraft and fly towards the heading bug. Rollattitude from the on-side AHRS will add with the error signal in the flightdirector processor to center the command cue when the proper bank anglehas been achieved.

As the aircraft approaches the selected heading, the heading error signalgets smaller in size and the roll attitude signal will now command the pilotto roll the aircraft to a wings level condition. With the aircraft flying theselected heading, the following conditions will exist:

• Heading select error is zero

• Flight director command bar is centered

• Control wheel is centered

• Aircraft is maintaining the selected heading.



With the autopilot engaged, the flight director processor will generate thecommands as stated above, but will now send them to the autopilot forautomatic flight path steering. On the PFD the flight director command barmay move a little out of center and then return. With the autopilotsatisfying the flight director steering command, the command cue will becentered.

Input data used by the heading select control law includes selectedheading, actual heading, TAS and roll attitude.

2 Low Bank Submode

The low bank submode allows pilot selection of reduced bank angle limitsfor the heading select mode. The mode is selected by pressing the lowbank button (BNK) on the GC-550 Guidance Control Unit. The bank anglelimit will be reduced from +27° to +14°. The mode will only be annunciatedwhile the heading select mode is active, but will remain selected and willreactivate and annunciate if heading select is made active again. The lowbank mode is canceled by pressing the BNK button while the annunciatoris illuminated.

Low Bank mode is automatically selected by climbing through 25,000 feet.Automatic canceling of Low Bank occurs descending through 24,750 feet.

3 VOR/VOR Approach Mode

When the VOR mode is armed, the flight director processor comparesactual aircraft heading against selected aircraft course, as determined bythe position of the course select pointer on the master PFD. Thedifference is the course error signal.

The lateral beam sensor (LBS) is computing when to capture the VORbeam. At VOR capture, the heading select mode is dropped and the flightdirector processor generates a command to bank the aircraft and getaligned on the VOR beam center.

With the autopilot not engaged, the VOR error signal is presented on thePFD flight director command bar as a computed steering command for thepilot to bank the aircraft and fly towards the course pointer. Roll attitudefrom the on-side AHRU will add with the error signal in the flight directorprocessor to center the command cue when the proper bank angle hasbeen achieved.



As the aircraft approaches the selected course, the course error signalgets smaller in size and the roll attitude signal will now command the pilotto roll the aircraft to a wings level condition. With the aircraft flying theselected course, the following conditions exist:

• Course select error is appropriate to maintain VOR beam center

• Radio deviation is zero

• Flight director command bar is centered


• Aircraft is tracking the selected VOR radial.

With the autopilot engaged, the flight director processor will generate thecommands as stated above, but will now send them to the autopilot forautomatic flight path steering. On the PFD, the flight director commandbar may move a little out of center and then return. With the autopilotsatisfying the flight director steering command, the flight director commandbar will be centered.

As the aircraft flies over the VOR station, the flight director processormonitors for entry into the zone of confusion above the VOR station. WithDME valid, when the aircraft is NAV on course (NOC) and DME =(barometric altitude/cosine of 30 degrees), the system will go into overstation sensing (OSS) and ignore the radio input.

With DME not valid, or not available, the system monitors beam deviationand beam rate for the OSS function. Beam deviation must be greater than75 mV and beam rate of change greater than 7.5 mV/sec. When radiodeviation drops below 75 mV, a 20 second clock is started (4 seconds inVAPP). At the end of this time, the radio input is again made part of theVOR equation. The time delay is to ensure the aircraft has cleared thezone of confusion.

The input data used by the VOR control law includes selected course, VORbearing, DME, TAS, baro corrected altitude and roll attitude.

4 Localizer/Back Course Modes

When the localizer mode is armed, the flight director processor comparesactual aircraft heading against selected aircraft course, as determined bythe position of the course select pointer on the coupled side PFD. Thedifference is the course error signal.

The lateral beam sensor (LBS) is computing when to capture the localizerbeam. At localizer capture, the heading select mode is dropped and theflight director processor generates a command to bank the aircraft and getaligned on the localizer beam center.



With the autopilot not engaged, the localizer error signal is presented onthe PFD flight director command bar as a computed steering command forthe pilot to bank the aircraft and fly towards the course pointer. Rollattitude from the AHRS will add with the error signal in the flight directorprocessor to center the command cue when the proper bank angle hasbeen achieved.

As the aircraft approaches the selected course, the localizer error signalgets smaller in magnitude and the roll attitude signal will now command thepilot to roll the aircraft to a wings level condition. With the aircraft trackingthe localizer beam, the following conditions exist:

• Course select error is appropriate to maintain localizer beam center.


• Command cue is centered


• Aircraft is tracking the localizer beam.

With the autopilot engaged, the flight director processor will generate thecommands as stated above, but will now send them to the autopilot forautomatic flight path steering. On the PFD, the command cue may move alittle out of center and then return. With the autopilot satisfying the flightdirector steering command, the command cue will be centered.

5 LNAV Mode

When the LNAV mode is active, the flight director processor in the IC-600IAC receives computed steering commands from the LNAV unit over anARINC 429 bus. These commands allow the flight director to fly the activeflight plan as displayed on the FMS control display unit (CDU).

On the PFD the course select pointer is now a desired track pointer and ispositioned automatically by the FMS.

The input data used by the LNAV control law includes the compositesteering command from the FMS unit and roll attitude from the AHRU.




7181

PRIMARY RADIOSYSTEM BUS

HL

88102

SECONDARY RADIOSYSTEM BUS

HL

LRN (FMS)

164J1A

164J1B

C190J2A-21C190J2A-22

AZ-850 MADC NO. 1

6061

9J1

C190J2A-43C190J2A-44

AA-300 RADIOALTIMETER NO. 1 W

N

20J1

PC-400 AUTOPILOTCONTROLLER

B

H

TURN KNOB OUTPUT

TURN KNOB OUTOF DETENT

129J1

70

73

75

1326

8384

5150

4344

2324

3952

2122

1920

5354

190J2B

190J2B

190J2A

190J2B

HL

HL

HL

HL

HL

HL

HL

HL

TCS INPUT

TK INPUT

TKOD INPUT

ARINC 429 INWINDSHEAR

ARINC 429 INSEC AHRS

RAD ALT +RAD ALT -

ARINC 429 INSEC MADC

ARINC 429 INPRI MADC

TK +15V OUTPUT

ARINC 429 OUTLRN/FMS

ARINC 429 INLRN/FMS



48FD1 ON PB

29GA DISCRETE IN

22HDG PB IN

24NAV PB IN25APR PB IN

36LAMP DRVR OUT FD1

21LAMP TEST IN

PILOT'S

COPILOT'S

GA SWITCHESON THROTTLES

SIGNALGND

LAMP TEST

C190J1B-21

190J1B

190J2B

5556

59

129J1-C

99

190J2A

106

104

1516

49

5758

64

C190J2A-57C190J2A-58

AIR

WOW

20J1-Y193RMP-2K193RBP-3CC190J2A-49

PILOT'S TCS

COPILOT'S TCS

C190J2B-70C115J1-32

AUTOPILOTSERVO

POWER

34

32

115J1

35

TCS INPUT CHASSIS GND 45

HL

DC/SG BUS

1

5DAY/NIGHT

SWITCH

6

3O-5VDC

EDGELTG

4DC POWER GND

68HDG PUSH TO SYNC6371HDG PUSH TO SYNCCRS1 PUSH TO SYNC6270CRS1 PUSH TO SYNC

4DC GND

1SIGNAL GND

#2 COMMON4914#2 SET KNOB4813HDG KNOB INPUT#2 SET KNOB4712

HDG

#1 COMMON4611#1 SET KNOB4510CRS1 KNOB INPUT#1 SET KNOB449

CRS1

DC-550 DISPLAYCONTROLER NO. 1

29

2625

21

33

98105 28

27

14131211

10

7

+15V MODE POWER

APR PUSH BUTTONNAV PUSH BUTTONBNK PUSH BUTTONHDG PUSH BUTTON

CPL PUSH BUTTON

FD1 PUSH BUTTON

GC-550 GUIDANCECONTROL UNITIC-600 IAC NO. 1

AD-51326-R1@

CTURN KNOB +15V INPUT 190J2B-59

C190J2B-26C190J2B-13

C1J1B-E5FIG. 2-10-6

C1J1B-E6

C9J1-60FIG. 2-10-6 C9J1-61

CPL PB INBANK PB IN

+15V EXTERNAL OUT

LAMP DRVR CPL OUTLAMP DRVR HDG OUT

LAMP DRVR BANK OUTLAMP DRVR NAV OUTLAMP DRVR APR OUT

+-

DC/SG BUS

RAD ALT VALID

WOW INPUT

HL

IC BUS

C190J2A-64

C190J1B-29

CPL LEFT LAMP INHDG LAMP INBNK LAMP INNAV LAMP INAPR LAMP IN

LIGHTING COMMON

5V PANEL LIGHTING

ANNUNCIATOR LTG

PILOT MODE GND

FD1 LAMP IN

11J1

C190J2B-75

C190J2A-50C190J2A-51

115J1

C115J1-12C115J1-13C115J1-14

C115J1-71

C190J2B-56C190J1B-24C190J1B-25

C190J1B-22

C190J2B-55

C190J2A-98C190J2A-105C190J2A-104

C190J2A-106

60 -15V EXTERNAL OUTTURN KNOB -15V INPUT A

WINDSHEAR COMPUTER

ARINC 429 BUS 1HL

AH-800 AHRS NO. 1G7G8

1J1B

2930

190J2A

HL

ARINC 429 IN PRI AHRSARINC 429 OUT 1

HL

E5E6ARINC 429 OUT 2

HL

C190J2A-83C190J2A-84

Figure 2-10-5. Flight Director Lateral Modes Interface - Pilot’s Side




7181


HL

88102


HL

LRN (FMS)

AH-800 AHRS NO. 2G7G8

164J1A

C164J1B

C1J1B

190J2A-21190J2A-22

AZ-850 MADC NO. 2

6061

C9J1

190J2A-43190J2A-44

AA-300 RADIOALTIMETER R/T W

N

20J1

WINDSHEAR COMPUTER


HTURN KNOB OUTOF DETENT

129J1

70

75

1326

8384

5150

4344

24

2930

3952

2122

1920

5354

C190J2B

C190J2A

C190J2B

190J2A

C190J2B

HL

HL

HL

HL

HL

HL

HL

HL

HL

TCS INPUT

TKOD INPUT



RAD ALT +RAD ALT -


ARINC 429CH 3 IN

ARINC 429CH 4 IN

ARINC 429 OUTLRN/FMS

ARINC 429 INLRN/FMS



48FD2 ON PB

29GA DISCRETE IN

22HDG PB IN

24NAV PB IN25APR PB IN

36LAMP DRVR OUT FD2

21LAMP TEST IN

PILOT'S

COPILOT'S


SIGNALGND

LAMP TEST

190J1B-21

C190J1B

C190J2B

5556

99

C190J2A

106

104

1516

49

5758

64

190J2A-57190J2A-58

AIR

WOW

20J1-Y193RMP-2K193RBP-3C190J2A-49

PILOT'S TCS

COPILOT'S TCS

C190J2B-70C115J1-32

AUTOPILOTSERVOPOWER

34

32

115J1

35


HL

DC/SG BUS

1

5DAY/NIGHT

SWITCH

6

3O-5VDC

EDGELTG

4DC POWER GND

68HDG PUSH TO SYNC6371HDG PUSH TO SYNCCRS2 PUSH TO SYNC6270CRS1 PUSH TO SYNC

4DC GND

1SIGNAL GND

#2 COMMON4914#2 SET KNOB4813HDG KNOB INPUT#2 SET KNOB4712

HDG

#5 COMMON4611#5 SET KNOB4510CRS2 KNOB INPUT#5 SET KNOB449

CRS2


29

2636

22

98105 28

27

14131211

10

20

APR PUSH BUTTONNAV PUSH BUTTONBNK PUSH BUTTONHDG PUSH BUTTON

CPL PUSH BUTTON

FD2 PUSH BUTTON


AD-51327-R1@

190J2B-26190J2B-13

1J1B-E5FIG. 2-10-5

1J1B-E6

9J1-60FIG. 2-10-5

9J1-61

CPL PB INBANK PB IN

LAMP DRVR CPL OUTLAMP DRVR HDG OUT

LAMP DRVR BANK OUTLAMP DRVR NAV OUTLAMP DRVR APR OUT

+-

DC/SG BUS

RAD ALT VALID

WOW INPUT

HL

IC BUS

190J2A-64

190J1B-29

CPL RIGHT LAMP INHDG LAMP INBNK LAMP INNAV LAMP INAPR LAMP IN

LIGHTING COMMON

5V PANEL LIGHTING

ANNUNCIATOR LTG

PILOT MODE GND

FD2 LAMP IN

11J1

190J2B-75

190J2A-50190J2A-51

C115J1

115J1-12115J1-13115J1-14

115J1-71

190J2B-56190J1B-24190J1B-25

190J1B-22

190J2B-55

190J2A-98190J2A-105190J2A-104

190J2A-106

ARINC 429 BUS 1HL

ARINC 429 OUT 1HL

23

E5E6ARINC 429 OUT 2

HL

190J2A-83190J2A-84

Figure 2-10-6. Flight Director Lateral Modes Interface - Copilot’s Side



For the descriptions and figures used in this section, the assumption is made that thepilot’s flight director is master.

(2) Heading Select (HDG) Mode

The heading select mode is used to intercept and maintain a magnetic headingreference. The mode is engaged by pressing the HDG button on the GC-550Guidance Control Unit, or by arming the LOC, BC, VOR, LNAV or VAPP modes.HDG will be annunciated on the PFD. Engaging the heading select mode will resetall previously selected active lateral modes.

The heading bug on the PFD is positioned around the compass card to the headingthe pilot desires to intercept, using the HDG knob on the GC-550 Guidance ControlUnit. The heading select error signal sent to the flight director processor is thedifference between the actual aircraft heading and the selected aircraft heading. Theflight director processor will now generate the proper roll command to intercept andmaintain the pilot-selected heading.

A normal heading select function is implemented based on an activeweight-on-wheels discrete (aircraft on ground). When in the normal heading selectmode, the flight director command bar steers in the direction in which the headingselect bug is the shortest distance from the fore lubber line on the PFD. If theheading select bug is exactly on the aft lubber line, the flight director steeringcommand may be initiated in either direction.

A smart heading select function is implemented based on an inactiveweight-on-wheels discrete (aircraft in air). When the smart heading select mode isactivated, the flight director steering command will be the shortest distance to thebug, or in the direction of travel of the bug from the fore lubber line.

Heading Select mode operating limits are listed in Table 2-10-5.

The heading select mode is canceled by:

• Capture of any other lateral steering mode

• Selecting go-around

• SG reversionary selection

• Pressing the HDG button on the GC-550 Guidance Control Unit

• Activation of the CPL pushbutton on the GC-550 Guidance Control Unit

• Turn knob out of detent with autopilot engaged.

Table 2-10-5. Heading Select Mode Operating Limits

Mode Parameter Value

Heading Select Roll Angle limitLow bank limit

± 27°± 14.0°

Roll Rate Limit 3.0°/sec



(3) Heading Select Mode Engage/Reset/Disengage Logic

(a) Engage Logic

Required valids:

• Flight director

• On-side AHRU

With the above conditions met, pressing either the:

• HDG button on the GC-550 Guidance Control Unit

• NAV button on the GC-550 Guidance Control Unit with either a VOR,LNAV or LOC NAV source tuned

• APR button on the GC-550 Guidance Control Unit with either a LOC orVOR frequency tuned.

will engage the heading select mode.

(b) Reset/Disengage Logic

Reset means that a condition has occurred that has canceled the mode, but itcan be re-engaged.

Disengage means that a condition has occurred that has canceled the modedue to a fault and the mode cannot be re-engaged until the fault is cleared.

The heading select mode will automatically be reset/disengaged, if any of thefollowing conditions happen:

• Pressing the HDG button on the GC-550 Guidance Control Unit (reset)

• Any lateral flight director mode captured (reset)

• Selecting go-around (reset)

• Changing the displayed heading source on the PFD (reset)

• Anytime the flight guidance system is powered up (reset)

• Activating the GC-550 CPL pushbutton (reset)

• Flight director system not valid (disengage)

• On-side AHRU not valid (disengage)


NOTE: If the flight director goes invalid, the command cuewill bias from view. If the AHRU goes invalid, themode clears, as well as attitude/heading failure flagsdisplayed on the EDS display.



(4) VOR (NAV) Mode

See Figures 2-10-7 through 2-10-13 and Table 2-10-6.

The VOR mode provides for automatic intercept, capture and tracking of a selectedinbound or outbound VOR radial, utilizing the selected VOR navigation sourcedisplayed on the on-side PFD. The navigation source displayed on the PFD is afunction of the NAV source buttons located on the DC-550 Display Controller. Priorto engaging the mode the pilot would perform the following:

• Tune the navigation receiver to the desired VOR frequency

• Select NAV as the navigation source on the DC-550 Display Controller

• Set the course pointer on the on-side PFD for the desired course to be flown

• Set the heading bug on the PFD to the desired intercept heading for the selectedcourse.

With the aircraft outside the normal capture range of the VOR signal (typically thecourse deviation on the PFD is greater than two dots), the pilot presses the NAVbutton on the GC-550 Guidance Control Unit. The HDG and NAV buttons on theGC-550 will illuminate. HDG in green and VOR ARM in white are also annunciatedon the PFD. The IC-600 IAC is now armed to capture the VOR signal and isgenerating a roll command to fly the heading select mode.

SELECTEDCOURSE090°

SELECTED RADIAL (270°)

SELECTEDVOR STATION

SELECTEDHEADING OR TRACK30°

TYPICAL CAPTURE POINTON HSI

AD-50969@

Figure 2-10-7. VOR ARM Pictorial



180

160

220

24010500

9500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

200 100 00

169 10000HDG ALT

E30.01 IN

030 CRS

VOR1

26.5

25

HDG030

VOR1

NM

KTSGSPD

280

M

20

80

-2000

AP

VOR

AD-50970@

YD

Figure 2-10-8. VOR (NAV) Mode Armed

When reaching the lateral beam sensor (LBS) trip point, the flight directorautomatically drops the heading select mode and switches to the VOR capturephase. The following is observed on the PFD:

• The white VOR annunciator extinguishes

• The green HDG annunciator extinguishes

• A green VOR is annunciated and is enclosed in a white box for 8 seconds toemphasize the capture phase of operation.

The IC-600 IAC now generates the proper roll command to bank the aircraft tocapture and track the selected VOR radial.



SELECTEDCOURSE090°

SELECTEDVOR STATION

TYPICAL CAPTURE POINTAD-50974@

Figure 2-10-9. VOR Capture Pictorial

When the course select pointer was set on the PFD using the appropriate courseknob on the GC-550 Guidance Control Unit, the course select error signal wasestablished. This signal represents the difference between actual aircraft headingand the desired aircraft course.

The radio signal is routed from the navigation receiver to the IC-600 IAC, where theradio signal is processed and lateral gain programmed.

Lateral gain programming is performed as a function of DME distance to the station(if available) and TAS. This gain programming adjusts for the aircraft either comingtoward or moving away from the VOR station.

NOTE: Avoid, if possible, operating in DME hold during VOR capture andtracking operation. When in DME hold the computer cannot useDME distance for gain programming.



180

160

220

24010500

950030

30

30

30

3

2

1

0

1

2

3

200 100 00

169 10000VOR ALT

E30.01 IN

090 CRS

VOR1

24.1

25

HDG030

VOR1

NM

MINTTG

5.0

YDM

20

80

-2000

AP

AD-50973@

Figure 2-10-10. VOR (NAV) Mode Capture

When flying a VOR intercept, the optimum intercept angle should be 45 degrees orless. If the intercept angle is greater than 45 degrees, then course cut limiting mayoccur.

The course cut limiter functions primarily when approaching the desired VOR radialat an intercept angle greater than 45 degrees and at a high rate of speed. Itsfunction is to limit steering commands to 45 degrees, which forces a flight path to geton the selected radial sooner to prevent overshooting the VOR beam center.Typically, the roll command will make an initial heading change, then level out andfly toward the beam, then make a second heading change to get lined up on thecenter of the selected radial.



SELECTEDCOURSE090°

SELECTEDVOR STATION

TYPICAL CAPTURE POINT

AD-51270@

SECONDHEADINGCHANGEINITIAL

HEADINGCHANGE

Figure 2-10-11. VOR Course Cut Limiting

When the aircraft satisfies VOR track conditions, the course error signal is removedfrom the lateral steering command. This leaves NAV on course (NOC) and DMEgain programming (if available) to track the VOR signal and to compensate for beamstandoff in the presence of a crosswind. The system will automatically compensatefor a crosswind of up to 45 degrees course error.



180

160

220

24010500

9500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

200 100 00

169 10000VOR ALT

E29.92 IN

090 CRS

VOR1

20.9

25

HDG030

ADF2

VOR1

NM

MINTTG

5.0

YDM

20

80

-2000

AP

AD-50975@

Figure 2-10-12. VOR Track

As the aircraft approaches the VOR station, it will enter a zone of unstable radiosignal. This zone of confusion radiates upward from the station in the shape of atruncated cone. In this area, the radio signal becomes highly erratic and it isdesirable to remove it from the roll command. The overstation sensor (OSS)monitors entry into the zone of confusion and removes radio deviation from the rollcommand. The system also uses the collocated DME signal (if available) to adjusttracking gains.



When over the VOR station (Figure 2-10-13), the system will accept and follow acourse change of up to 90 degrees.

The navigation (VOR) mode is canceled by:

• Pushing the NAV button on the GC-550 Guidance Control Unit


• Selecting another lateral mode active

• NAV source change (radio frequency on the selected NAV receiver.)

• Displayed heading invalid

• On-side attitude invalid

• Displayed NAV source invalid

• On-side air data invalid

• SG reversion

• Activation of the GC-550 CPL pushbutton


AD-50976@

ZONE OF CONFUSION

OSS A0SS1 A0SS2

Figure 2-10-13. VOR Overstation



(5) VOR Approach (VAPP) Mode

A VOR approach is completed in a manner similar to normal VOR. However, insteadof selecting the mode with the NAV button on the GC-550 Guidance Control Unit, theAPR button is pushed. The APR button annunciator illuminates and VAPP isdisplayed in white on the PFD. The flight director now applies the gains appropriatefor an approach. Upon capture of the selected course, the PFD displays VAPP ingreen. Refer to Table 2-10-6 for system operating limits.

Table 2-10-6. VOR/VOR Approach Operating Limits


VOR orVAPP

Capture:

Beam Intercept Angle Up to 90°

Capture Point Function of DME, beamdeviation, beam closure rateand course error.MIN Trip Point: ± 30 mV dcMAX Trip Point: ± 175 mV dc

Roll Angle Limit ± 27.5°

Roll Rate Limit 3.0°/sec VOR5.5°/sec VAPP

Course Cut Limit 45° during capture

VOR NOC Roll Angle Limit ± 17°

Roll rate Limit 1.0°/sec VOR5.5°/sec VAPP

Crosswind Correction Up to 45° course error VORUp to 30° course error VAPP

Over Station:

Course Change Up to 90°

Roll Angle LimitRoll Rate Limit

± 17°3°/sec



(6) VOR/VAPP Engage/Reset/Disengage Logic

(a) VOR ARM Engage Logic

Required valids:

• Flight Director

• On-side AHRU

• On-side MADC

• Displayed NAV source.

With the above conditions met, and:

• VOR selected as the NAV source on the DC-550 Display Controller,pressing the NAV button on the GC-550 Guidance Control Unit will arm theVOR mode.

(b) VOR Arm Reset Logic



VOR arm will automatically be reset if any of the following conditions happen:

• VOR is captured


• NAV source change

• Any other lateral flight director mode active

• Pressing the NAV button on the GC-550 Guidance Control Unit

• Changing the displayed heading source

• Anytime the flight director system is powered up





(c) VOR Arm Disengage Logic

The VOR arm mode will automatically disengage, if any of the followingconditions happen:

• Flight Director not valid

• On-side AHRU not valid

• Tune to a localizer frequency

• Displayed NAV source not valid

• SG reversionary switching

• On-side MADC not valid.

NOTE: Loss of NAV valid causes the flight director commandbar to go out of view, while maintaining modeannunciation and heading hold command. If the NAVremains invalid after 5 seconds, the mode will becanceled. If NAV valid returns within 5 seconds, thecommand bar will re-appear.

(d) VOR Capture Engage Logic

Required valids:

• Flight Director

• On-side AHRU

• Displayed NAV source

• On-side MADC.

With the above conditions met and the VOR mode is armed plus 1 second andthe LBS trips, the VOR mode will automatically transition from arm to capture.

(e) VOR Capture Reset Logic

VOR capture will automatically reset, if any of the following conditions happen:


• Any other lateral mode active


• Pressing the APR button on the GC-550 Guidance Control Unit








(f) VOR Capture Disengage Logic

The VOR mode will automatically disengage, if any of the following conditionshappen:




• NAV source not valid

• On-side MADC not valid



(g) NAV On Course (NOC) Logic

For VOR, NOC is defined as follows:

• LBS has tripped

• Course error less than 22°

• Bank angle less than 6°

• Beam deviation less than 75 mV.

When all the above conditions have existed simultaneously for 8 seconds,NOC is latched. As a function of NOC latching, the flight director processorstarts cross wind correction.

NOTE: Loss of NAV valid will cause the flight directorcommand bar to bias out of view. If the NAV remainsinvalid after 5 seconds, the mode will be canceled.

(h) VOR OSS Engage Logic

Required valids:

• Flight Director

• On-side AHRU


• On-side MADC.



With the above conditions met, and VOR is NOC, the flight director processormonitors for:

• Beam deviation greater than 75 mV

• Beam rate of change greater than 7.5 mV/second or with DME valid andnot Hold (barometric altitude/cosine of 30°) = DME.

If either of these conditions exist at the same time, the flight director processorassumes the aircraft is overstation and inhibits the radio input.

(i) VOR OSS Reset Logic









(j) VOR OSS Disengage Logic








The logic for the VOR Approach mode is identical to the logic for the VORmode except for the following:

• Pressing the APR button instead of the NAV button on the GC-550Guidance Control Unit.



(7) Localizer (NAV) and Back Course (BC) Modes

See Figures 2-10-14 through 2-10-22 and Table 2-10-7.

The localizer mode provides for automatic intercept, capture, and tracking of thefront course localizer beam to line up on the centerline of the runway in use. Theback course localizer mode allows the pilot to fly a back course localizer intercept.Prior to mode engagement, the pilot would perform the following:

• Tune the navigation receiver to the published front course localizer frequency forthe runway in use

• Press the NAV button on the DC-550 Display Controller to select LOC as thenavigation source

• Set the course pointer on the coupled side PFD for the inbound runway heading

• Set the heading bug on the coupled side PFD for the desired heading to performa course intercept.

The PFD now displays the relative position of the aircraft to the center of thelocalizer beam and the desired inbound course. With the heading bug set for courseintercept, the heading select mode is automatically used to perform the intercept.Outside the normal capture range of the localizer signal, pressing the NAV button onthe GC-550 Guidance Control Unit will cause the PFD to annunciate HDG in greenand LOC in white.

The aircraft is now flying the desired heading intercept and the flight director isarmed for automatic localizer beam capture.

090°INBOUNDCOURSE

LOCALIZERTX



AD-50978@

RUNWAY

Figure 2-10-14. Localizer ARM Pictorial



160

140

180

5000

4000

2002020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

175 45 00

169 4500HDG ALT

E

29.85 IN

090 CRS

ILS1

13.2

HDG040

ADF

NM

KTSGSPD

150

YDM

2350

20

80

-2000

AP

LOC

AD-50980@

350 RA

Figure 2-10-15. Localizer (NAV) Mode ARM

With the aircraft approaching the selected course intercept, the LBS is monitoringlocalizer beam deviation, beam rate, and TAS. At the computed time, the LBS willtrip and capture the localizer signal. The flight director processor now drops theheading select mode and generates the proper roll command to bank the aircrafttoward localizer beam center. When the LBS trips, the following is observed on thePFD:

• The green HDG annunciation extinguishes

• The white LOC annunciation extinguishes

• The green LOC annunciation comes on.



090°INBOUNDCOURSE

LOCALIZERTX

AD-50979@

RUNWAY


Figure 2-10-16. Localizer Capture Pictorial

The IC-600 IAC flight director processor now generates the proper roll command tobank the aircraft to capture and track the selected localizer signal.

NOTE: When flying a localizer intercept, the optimum intercept angle is45 degrees or less. If the intercept angle is greater than 45degrees, course cut limiting can occur as previously described inthe VOR mode of operation.

When the course select pointer was set on the coupled side PFD using theappropriate course knob on the GC-550 Guidance Control Unit, the course selecterror signal was established. This signal represents the difference between actualaircraft heading and selected aircraft course.

Lateral gain programming is required to adjust the gain applied to the localizer signaldue to the aircraft approaching the localizer antenna and beam convergence causedby the directional properties of the localizer antenna. The lateral gain programmer iscontrolled by the change in radio altitude when the aircraft is below 2400 ft radioaltitude and the radio altimeter is valid. If the radio altimeter is not valid, then gainprogramming occurs as a function of localizer beam capture.



The localizer mode is canceled by:

• Pushing the NAV button on the GC-550 Guidance Control Unit



• Changing navigation sources

• Displayed NAV source invalid

• On-side AHRU invalid

• On-side MADC invalid


• Activation of the CPL pushbutton


A Back Course Localizer Approach is automatically selected by the IAC as a functionof the course error (aircraft heading versus runway heading), during the arm phaseof the mode. If the course error is < 105 degrees LOC is selected as the arm mode.If the course error is > 103 degrees BC is selected as the arm mode.

The ILS approach mode is set up and flown identically to the localizer mode with thefollowing differences:

• On the GC-550 Guidance Control Unit, the APR button is pressed instead of theNAV button. This will arm both the localizer and glideslope modes for automaticcapture to fly a fully coupled ILS approach.

• The system is interlocked such that glideslope capture is inhibited unlesslocalizer capture occurs first.

When the APR button is pressed on the GC-550 Guidance Control Unit, the systemis now set to automatically capture the localizer and glideslope signals. On the PFDthe following annunciators would illuminate:

• HDG in green

• LOC in white

• GS in white.

The system is interlocked so that glideslope capture is inhibited unless localizer iscapture or track. Glideslope capture cannot occur when flying a back courselocalizer approach. Input data used by the localizer control law includes selectedcourse, localizer deviation, TAS, radio altitude, middle marker and roll attitude.



140

120

160

180

4000

3000

AD-50981@

20

2030

30

30

30

3

2

1

0

1

2

3

150 35 00

169 3500LOC ALT

E

29.85 IN350 RA

090 CRS

ILS1

8.5

HDG360

ADF

NM

KTSGSPD

145

YDM

2150

20

40

-2000

AP

RF

Figure 2-10-17. Localizer Capture



120

100

160

1802000

1500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

140 1860

169 1500LOC VS

E

29.85 IN350 RA

090 CRS

ILS1

13.1

HDG060

ADF1

NM

KTSGSPD

110

M

-1100

640

40

-1000

ASEL

AP

RF

AD-50983@

Figure 2-10-18. Localizer Track



090°INBOUNDCOURSE

LOCALIZERTX

AD-50985@

RUNWAY

340°

TYPICALCAPTURE

POINT

Figure 2-10-19. Back Course Mode Pictorial



140

120

180

2005000

4000

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

160 45 00

169 4500HDG ALT

E

30.02 IN450 RA

090 CRS

ILS1

13.1

HDG320

ADF1

NM

KTSGSPD

150

M

2130

20

80

-2000

BC

AD-50984@

Figure 2-10-20. Back Course Arm



140

120

180

2004500

3500

30

30

30

30

3

2

1

0

1

2

3

160 40 00

169 4000BC ALT

R

29.92 IN450 RA

090 CRS

ILS1

10.5

25

HDG320

VOR2

ADF1

NM

KTSGSPD

145

YDM

2130

20

80

-2000

AP

E

AD50986@

Figure 2-10-21. Back Course Capture



100

80

140

1602500

1500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

120 20 00

169 1600BC VS

AP

30.02 IN450 RA

090 CRS

ILS1

3.6

HDG320

ADF1

NM

KTSGSPD

110

YDM

-1100

850

20

80

-1100

AP

ASEL

R

E

AD-50987@

Figure 2-10-22. Back Course Track



Table 2-10-7. Localizer (LOC) and Back Course (BC) Mode OperatingLimits


LOC/BC LOC or BC Capture:

Beam Intercept AngleUp to 90°

Capture Point Function of beam deviation, beamclosure rate and localizer courseerror

MIN Trip Point: ± 60 mV dcMAX Trip Point: ± 175 mV dc

Roll Angle Limit ± 27.5°


Course Cut Limit 30° during capture

LOC or BC NOC:

Roll Angle Limit 17°

Roll rate Limit 5.5°/sec

Crosswind Correction Up to 30° course error

Gain programming Function of radio altitude

(8) Localizer/Back Course Mode Engage/Reset/Disengage Logic

(a) LOC/BC ARM Engage Logic

Required valids:

• Flight Director

• On-side AHRU

• On-side MADC

• Displayed NAV Source.


• NAV selected as the NAV source on the DC-550 Display Controller and thesource tuned to an ILS frequency.

Pressing the NAV button on the GC-550 Guidance Control Unit will arm theLOC/BC mode. Pressing the APR button on the GC-550 Guidance ControlUnit will arm the LOC and GS modes.



(b) LOC/BC Arm Reset Logic



LOC/BC arm will automatically be reset if any of the following conditionshappen:

• LOC/BC is captured





• Anytime the flight guidance system is powered up



(c) LOC/BC Arm Disengage Logic

The LOC/BC arm mode will automatically disengage if any of the followingconditions happen:

• Flight Guidance not valid


• Not tuned to a localizer frequency





NOTE: If the flight director system or the on-side AHRS goinvalid,the mode clears as well as the EDS display(attitude fail). If the NAV sensor fails, the mode willstay engaged, but the flight director command bar willbias out of view.



(d) LOC/BC Capture Engage Logic

Required valids:

• Flight Director

• On-side AHRU


• On-side MADC.


• The LOC/BC mode is armed plus 1 second and

• The LBS trips or deviation less than 35 mV.

The LOC/BC mode will automatically transition from arm to capture.

(e) LOC/BC Capture Reset Logic

LOC/BC capture will automatically reset if any of the following conditionshappen:



• Pressing the NAV or APR button on the GC-550 Guidance Control Unit



• Anytime the flight guidance system is powered up





(f) LOC/BC Capture Disengage Logic

The LOC/BC mode will automatically disengage if any of the followingconditions happen:



• Not tuned to a localizer frequency

• Displayed NAV source not valid plus 5 seconds




NOTE: Loss of NAV valid will cause the flight directorcommand bar to bias out of view while maintainingmode annunciation. If the NAV remains invalid after5 seconds, the mode will be canceled.

(9) Long Range Navigation

See Figures 2-10-23 and 2-10-24 and Table 2-10-8.

Should the pilot select the long range navigation (FMS) source on the DC-550Display Controller, the NAV mode is flown similar to the VOR mode. Depending onthe installed FMS, the PRIMUS® 1000 DFGS can provide a variety of capabilities.These range as follows:

• Automatic arm and capture of LRN supplied track

• Immediate capture of the LRN supplied track.

Some common characteristics include:

• Instead of using course error and radio deviation, a composite lateral steeringcommand is utilized from the long range navigation computer by the IC-600 IAC

• The symbol generator function supplies the flight director with the requiredsteering commands

• This lateral steering command is gain programmed in the FMS and therefore isnot gain programmed again in the IC-600 IAC.



(a) Automatic FMS Arm/Capture

Pressing the NAV button on the GC-550 Guidance Control Unit will cause theHDG mode to annunciate ON. LNAV will be armed. At the proper point, theflight director will capture the FMS supplied track and LNAV will beannunciated in green on the PFD.

(b) Immediate FMS Capture

Pressing the GC-550 Guidance Control Unit NAV button will cause the flightdirector LNAV mode to annunciate in green on the PFD. The flight directorfunction will now be captured and direct the aircraft to a track intercept. Oncethe cross track deviation is centered, the desired track will be achieved.

AD-51330@

DESIRED TRACK

DESIRED TRACK

CAPTURE

CAPTURE

Figure 2-10-23. Long Range Navigation Capture Pictorial and Tracking



200

180

220

240

31500

30500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

210 310 20

169 31000FMS ALT

29.92 IN

349 DTK

FMS

79.1

125

HDG050

NM

KTSGSPD

365

YD

.630 M

M

40

00

-2000

AP

AD-50977@

Figure 2-10-24. Long Range Navigation Tracking

The LNAV mode is canceled by:



• Selecting another navigation source on the DC-550 Display Controller

• Selecting another heading source

• Selecting another lateral mode active






Table 2-10-8. LNAV Mode Operating Limits


LNAV Capture:

Beam Intercept AngleUp to 90°

Capture Point Function of groundspeed and theangular difference between actualand desired track.

Roll Angle Limit ± 32°


(10) LNAV Mode Engage/Reset/Disengage Logic

(a) LNAV Capture Logic

Required valids:

• Flight Director

• On-side AHRU

• On-side MADC

• FMS.

With the above conditions met and:

• The LNAV mode is armed and the aircraft is at the computed desired trackintercept point.

(b) LNAV Capture Reset Logic


LNAV capture will automatically reset if any of the following conditions happen:











(c) LNAV Capture Disengage Logic


The LNAV mode will automatically disengage if any of the following conditionshappen:



• Selected FMS source not valid






C. Flight Director Vertical (Pitch) Channel Functional Operation

(1) Flight Director Vertical (Pitch) Modes Interface

Figure 2-10-25 shows LRU interface for pilots side flight director vertical modes.Figure 2-10-26 shows LRU interface for copilots side flight director vertical modes.

The function of each LRU for each vertical mode is discussed in the followingparagraphs:

(a) AH-800 Attitude Heading Reference Unit (AHRU)

For all flight director vertical modes, the on-side AHRU provides actual aircraftpitch attitude, and rate of change terms for the on-side flight director. In theSpeed (SPD) and Flight Level Change (FLC) modes, the on-side AHRS alsoprovides a normal acceleration term. A longitudinal acceleration term is alsoprovided for the FLC mode. The normal acceleration term is also used in theAltitude Hold (ALT), Altitude Preselect (ASEL), and Glideslope (GS) modes.

The cross-side AHRU provides the same terms to the IC-600 IAC, but thesesignals are used for EDS/flight director and autopilot monitoring purposes only.

(b) AZ-850 Micro Air Data Computer (MADC)

The on-side AZ-850 MADC provides the on-side IC-600 IAC with an ARINC429 input of air data values including true airspeed (TAS). The TAS signal isused in some vertical flight director modes for gain programming. Theresponse of the aircraft should feel the same regardless of the aircraft’sairspeed and altitude. Since it requires less flight control surface deflection athigh speed and high altitude to complete a maneuver then it does at lowspeed and low altitude, changing the size of the signal as a function of TASachieves the desired results.

Should the AZ-850 MADC become invalid, a fixed bias TAS of 150 knots isused in the IC-600 IAC. The default value of TAS is set for the approachspeed region of flight.

Additionally, the AZ-850 MADC provides the on-side flight director with thefollowing vertical mode references:


• Indicated airspeed/Mach

• Altitude rate (vertical speed).

(c) GC-550 Guidance Control Unit

The GC-550 Guidance Control Unit provides the means for the pilot toengage/disengage all vertical flight director modes, with the exception ofGo-Around. The GC-550 also provides the controls to set the speed referencefor the SPD and VS modes, as well as setting the selected Altitude Preselectreference.



The GC-550 provides button input data directly to the IC-600 IAC. Speed setand selected altitude data is transmitted to both DC-550 Display Controllers.The DC-550 Display Controllers transmit the speed set and selected altitudedata to both IC-600 IACs.

(d) DC-550 Display Controller

The DC-550 Display Controller provides an RS-422 digital bus interface(DC/SG Bus) between itself and the IC-600 IAC. Speed set and selectedaltitude inputs are routed through the DC-550 to be put on the digital businterface to the IC-600 IAC.

(e) RT-300 Radio Altimeter

The radio altimeter provides an analog output of absolute altitude above theterrain. This signal is used by the flight director to gain program theglideslope signal. Gain programming is required due to the directionalqualities and beam convergence characteristics of the glideslope antenna.

As the aircraft approaches the runway, the glideslope signal appears to getstronger and the beam appears to get narrower. By reducing the gain on thesignal as a function of the change in radio altitude, the computed steeringcommand will not take the aircraft out of the glideslope beam envelope.

Should the radio altimeter be invalid, gain programming will start as a functionof glideslope capture and run down as a function of TAS and time. At themiddle marker gain programming will be synchronized to a preset value.

(f) RNZ-851 Integrated Navigation Unit

The RNZ-851 navigation unit provides an RSB output of glideslope deviationdata, as well as marker beacon data.

(g) IC-600 Integrated Avionics Computer (IAC)

The IC-600 IAC performs the following, as a function of which vertical mode isactive.

1 Pitch Attitude Hold

When only a lateral flight director mode is active, the IC-600 IAC willmemorize the pitch attitude of the aircraft at the time the lateral mode wasselected. This becomes the pitch attitude reference displayed on the PFD.

Pitch attitude can be changed by pushing and holding the TCS button andmaneuvering the aircraft to a new position. Releasing the TCS button willcause the IC-600 IAC to memorize this new attitude reference.



2 Vertical Speed Hold

The IC-600 IAC receives vertical speed information from the AZ-850MADC. This vertical speed information becomes the reference verticalspeed in the IC-600 IAC when the mode is engaged.

The vertical speed reference is changed as a function of turning the SPDknob on the GC-550 Guidance Control Unit. When the vertical speedmode is engaged, the speed set bug is displayed on the vertical speedscale, and the vertical speed air data command reference is displayedabove the vertical speed scale on the PFD.

a Flight Director Only

When the vertical speed mode is engaged and the autopilot is off,vertical speed steering commands are presented to the pilot on thePFD command bar. By flying the miniature aircraft symbol to the flightdirector command bar, the pilot will satisfy the flight director command.

b Autopilot Engaged

With the autopilot on and the vertical speed mode engaged, verticalspeed steering commands are sent to the autopilot for automatic flightpath steering.

3 Speed Hold

a Air Data

The IC-600 IAC receives IAS/Mach data from the AZ-850 MADC. Thisdata becomes the SPD reference in the IC-600 IAC when the mode isengaged.

The SPD reference is changed by turning the SPD set knob on theGC-550 Guidance Control Unit. When the SPD mode is engaged, themode synchronizes to the existing aircraft speed. The speedreference is shown as a digital display at the top of the airspeed tapeon the PFD. A speed set target bug is also displayed on the airspeedtape.

Input data used by the control law includes selected IAS/Mach, verticalspeed, actual IAS/Mach, longitudinal and normal acceleration terms.

b Flight Director Only

When the SPD mode is engaged and the autopilot is off, SPD steeringcommands are presented to the pilot on the PFD flight directorcommand bar. By flying the miniature aircraft symbol to the commandbar, the pilot will satisfy the flight director command.



c Autopilot Engaged

With the autopilot on and the SPD mode engaged, selected IAS/Machsteering commands are sent to the autopilot for automatic flight pathsteering.

4 Flight Level Change Mode

a Air Data

The IC-600 IAC receives IAS/Mach and vertical speed data from theAZ-850 MADC. When the FLC mode is engaged, as a function ofwhich profile (climb/descent) has been selected, actual IAS/Mach orvertical speed is compared against the profile programmed in theIC-600 IAC.

When the FLC mode is engaged in a climb, the applicable speedprofile target bug is displayed on the airspeed scale, and the selectedtarget airspeed profile is displayed above the airspeed scale on thePFD.

When the FLC mode is engaged in a descent, the applicable verticalspeed target is selected and displayed above the vertical speed scaleon the PFD.

Input data used by the control law includes profile airspeed or Machtarget, vertical speed, actual IAS/Mach, longitudinal and normalacceleration terms.

b Flight Director Only

When the FLC mode is engaged and the autopilot is off, FLC steeringcommands are presented to the pilot on the PFD flight directorcommand bar. By flying the miniature aircraft symbol to the commandbar, the pilot will satisfy the flight director command.

Since the reference speed is a fixed profile in software, it cannot bechanged with the SPD knob or TCS.

c Autopilot Engaged

With the autopilot on and the FLC mode engaged, the selectedIAS/Mach or vertical speed profile steering command is sent to theautopilot for automatic flight path steering.

5 Altitude Preselect Mode

The IC-600 IAC receives inputs of uncorrected pressure altitude, barocorrected pressure altitude and pilot selected altitude. The flight directorprocessor combines the pressure altitude inputs through complimentaryfiltering to obtain more precise barometric altitude data.In the IC-600 IAC the difference between actual aircraft altitude and



selected aircraft altitude (ASEL) is defined as the altitude error signal.

The altitude error signal is converted into a computed vertical speed signal.If the aircraft’s actual vertical speed is less than the computed verticalspeed, then ASEL remains armed. When the aircraft’s actual verticalspeed is greater than the computed vertical speed, the flight directorprocessor will capture the selected altitude and command the flaremaneuver.

There is no button to select the ASEL mode on the GC-550 GuidanceControl Unit. ASEL will automatically arm when the following conditionsexist simultaneously:

• Decreasing altitude error

• Computed vertical speed is greater than actual vertical speed

• Not in altitude hold

• Glideslope is not in capture or track mode

• Vertical speed is greater than 100 FPM for 3 seconds

• The target altitude is at least 250 feet from present altitude.

The input data used by the altitude preselect control law includes selectedaltitude, baro corrected altitude, vertical speed, and normal acceleration.


With the ASEL mode armed, the PFD will present a vertical steeringcommand of whichever other vertical mode is in use. When ASELtransitions from arm to capture, a vertical steering command ispresented to the pilot to flare the aircraft onto the selected altitude.

b Autopilot Engaged

With the autopilot engaged and the ASEL mode captured, steeringcommands are sent to the autopilot for automatic altitude capture.

6 Altitude Hold Mode

The IC-600 IAC receives an input of baro corrected pressure altitude fromthe AZ-850 MADC. The flight director processor compares actual aircraftaltitude against the altitude hold reference to generate the altitude holderror signal.

Input data used by the altitude hold control law includes pressure altitude,baro correction data, vertical speed, normal acceleration and pitch attitude.


With the ALT hold mode engaged, the PFD will present a verticalsteering command to the pilot to fly the aircraft back to the referencealtitude.



b Autopilot Engaged

With the autopilot on and the ALT hold mode engaged, the verticalsteering command is sent to the autopilot for automatic flightpathsteering.

7 ILS Approach Mode

When the ILS Approach mode (APR) is armed, the flight director processorlooks at TAS, vertical speed and glideslope deviation to determine when tocapture the glideslope. This is accomplished by the vertical beam sensor(VBS). When the glideslope is captured, the flight director processor willautomatically drop whichever other vertical mode was engaged and starttracking the glideslope signal. Should the localizer signal be lost for anyreason, the APR mode will also be dropped.

With the autopilot not engaged, the glideslope error signal is presented onthe PFD command bar as a computed steering command for the pilot toraise or lower the aircraft nose and fly back to glideslope beam center.Pitch attitude from the AHRS will add with the glideslope signal in the flightdirector processor to center the command bar when the proper pitchattitude has been achieved.

As the aircraft approaches the glideslope beam, the glideslope signal getssmaller in size and the pitch attitude signal will now command the pilot toreturn the aircraft to its landing attitude. With the aircraft tracking theglideslope beam, the following conditions exist:


• Command cue is centered


• Aircraft is tracking the glideslope beam.

With the autopilot engaged, the flight director processor generates thecommands stated above and sends them to the autopilot for automaticflight path steering. On the PFD, the command bar may move a little outof center and then return. With the autopilot satisfying the flight directorsteering command, the command bar will be centered.

8 Go-Around Mode

The GA mode is normally used to transition from an approach to land to aclimb out condition in the event of a missed approach. The pilot selectsGA mode by pushing the GA button located on either outboard throttlehandle. With GA mode selected, all flight director modes are canceled.




7181


HL

88102


HL

164J1A

164J1B

AZ-850 MADC NO. 1

6061

9J1

C190J2A-43C190J2A-44


N

20J1

PC-400 AUTOPILOTCOMPUTER

U

Y

129J1

70

68

69

1326

8384

5150

4344

2324

2930

1920

5354

190J2B

190J2A

190J2B

190J2A

190J2B

HL

HL

HL

HL

HL

HL

HL

TCS INPUT



RAD ALT +RAD ALT -



ARINC 429 INPRI AHRS



48FD1 ON PB

29GA DISCRETE

26ALT PB IN27SPD PB IN28FLC PB IN

30VS PB IN37LAMP DRVR OUT FLC36LAMP DRVR OUT FD1

21LAMP TEST IN

PILOT'S

COPILOT'S


SIGNALGND

LAMP TEST

C190J1B-21

190J1B

190J2B

55

59

129J1-C

99

190J2A

1031021011516

49

5758

64

C190J2A-57C190J2A-58

AIR

WOW

20J1-Y193RMP-2K193RBP-3CC190J2A-49

PILOT'S TCS

COPILOT'S TCS

C190J2B-70C115J1-32

AUTOPILOTSERVOPOWER

34

32

115J1

35


HL

DC/SG BUS

1

5DAY/NIGHT

SWITCH

6

3O-5VDC

EDGELTG

4

DC POWER GND

IAS/MACH SEL (SPD)6568IAS/MACH PUSH TO SYNC

4DC GND

1SIGNAL GND

#4 COMMON5520#4 SET KNOB5419SPD KNOB INPUT#4 SET KNOB5318

SPD

#3 COMMON5217#3 SET KNOB5116ALT KNOB INPUT#3 SET KNOB5015

ALT


343130

25

21

333219

171615

10

7

+15V MODE POWER

VS PUSH BUTTON

FLC PUSH BUTTONSPD PUSH BUTTONALT PUSH BUTTON

CPL PUSH BUTTON

FD1 PUSH BUTTON


AD-51328-R1@

C190J2A-50C190J2A-51

C190J2B-13C190J2B-26

+

-

PITCH WHEEL OUT+

-PITCH WHEEL INPUT

CPL PB IN

+15V EXTERNAL OUT

LAMP DRIVER OUT CPL

LAMP DRIVER OUT ALTLAMP DRIVER OUT IASLAMP DRIVER OUT VS

HLDC/SG BUS

HLIC BUS

RAD ALT VALID

WOW IN

94TO AUDIOWARNINGSYSTEM

ALT ALERTHORN OUT

C190J2A-64

LTG COMMON

5V PANEL LIGHTING

ANNUNCIATOR LTG

PILOT MODE GND

115J1

VS LAMP INSPD LAMP INALT LAMP IN

CPL LEFT LAMP IN

FD1 LAMP INFLC LAMP IN

11J1

C190J1B-29

C1J1B-E5FIG. 2-10-26

C1J1B-E6

C9J1-60FIG. 2-10-26 C9J1-61

WINDSHEAR COMPUTER

ARINC 429 BUS 1HL

AH-800 AHRS NO. 1 G7G8

1J1B

ARINC 429 OUT 1HL

E5E6ARINC 429 OUT 2

HL

C190J2A-83C190J2A-84

C190J1B-27C190J1B-28C190J1B-30

C190J1B-26

C190J2B-55

C190J1B-37

C190J2A-102C190J2A-101

C190J2A-103

C115J1-18C115J1-19C115J1-20

C115J1-15C115J1-16C115J1-17

C115J1-68

Figure 2-10-25. Flight Director Vertical Modes Interface - Pilot’s Side




7181


HL

88102


HL

C164J1A

C164J1B

AZ-850 MADC NO. 2

6061

C9J1

190J2A-43190J2A-44


N

20J1

PC-400 AUTOPILOTCOMPUTER

U

Y

129J1

70

68

69

1326

8384

5150

4344

2324

2930

1920

5354

C190J2B

C190J2A

C190J2B

C190J2A

C190J2B

HL

HL

HL

HL

HL

HL

HL

TCS INPUT



RAD ALT +RAD ALT -



ARINC 429 INPRI AHRS



48FD2 ON PB

29GA DISCRETE

26ALT PB IN27SPD PB IN28FLC PB IN

30VS PB IN37LAMP DRVR OUT FLC36LAMP DRVR OUT FD2

21LAMP TEST IN

PILOT'S

COPILOT'S


SIGNALGND

LAMP TEST

190J1B-21

C190J1B

C190J2B

55

99

C190J2A

1031021011516

49

5758

64

190J2A-57190J2A-58

AIR

WOW

20J1-Y193RMP-2K193RBP-3C190J2A-49

PILOT'S TCS

COPILOT'S TCS

C190J2B-70C115J1-32

AUTOPILOTSERVOPOWER

34

32

C115J1

35


HL

DC/SG BUS

1

5DAY/NIGHT

SWITCH

6

3O-5VDC

EDGELTG

4

DC POWER GND

IAS/MACH SEL (SPD)6568IAS/MACH PUSH TO SYNC

4DC GND

1SIGNAL GND

#4 COMMON5520#4 SET KNOB5419SPD KNOB INPUT#4 SET KNOB5318

SPD

#3 COMMON5217#3 SET KNOB5116ALT KNOB INPUT#3 SET KNOB5015

ALT


343130

36

223219

171615

10

20

VS PUSH BUTTON

FLC PUSH BUTTONSPD PUSH BUTTONALT PUSH BUTTON

CPL PUSH BUTTON

FD2 PUSH BUTTON


AD-51329-R1@

190J2A-50190J2A-51

190J2B-13190J2B-26

+

-

PITCH WHEEL OUT+

-PITCH WHEEL INPUT

CPL PB IN

LAMP DRIVER OUT CPL

LAMP DRIVER OUT ALTLAMP DRIVER OUT IASLAMP DRIVER OUT VS

HLDC/SG BUS

HLIC BUS

RAD ALT VALID

WOW IN

94TO AUDIOWARNINGSYSTEM

ALT ALERTHORN OUT

190J2A-64

LTG COMMON

5V PANEL LIGHTING

ANNUNCIATOR LTG

PILOT MODE GND

C115J1

VS LAMP INSPD LAMP INALT LAMP IN

CPL RIGHT LAMP IN

FD2 LAMP INFLC LAMP IN

11J1

190J1B-29

1J1B-E5FIG. 2-10-25

1J1B-E6

9J1-60FIG. 2-10-25 9J1-61

WINDSHEAR COMPUTER

ARINC 429 BUS 1HL

AH-800 AHRS NO. 2 G7G8

C1J1B

ARINC 429 OUT 1HL

E5E6ARINC 429 OUT 2

HL

190J2A-83190J2A-84

190J1B-27190J1B-28190J1B-30

190J1B-26

190J2B-55

190J1B-37

190J2A-102190J2A-101

190J2A-103

115J1-18115J1-19115J1-20

115J1-15115J1-16115J1-17

115J1-68

Figure 2-10-26. Flight Director Vertical Modes Interface - Copilot’s Side



The description and figures in this section assume that the pilots flight director is master.

(2) Pitch Attitude Hold

Refer to Table 2-10-9.

Pitch attitude hold is the basic vertical flight director mode. It is activated when aflight director lateral (roll) mode is selected without an accompanying vertical (pitch)mode. The PITCH hold mode is only active when the autopilot is engaged. There isa "PIT" annunciation for pitch attitude hold on the PFD in the vertical flight directormode field.

The pitch command on the PFD provides the pilot with a reference corresponding tothe pitch attitude existing at the moment the lateral flight director mode was selected.This pitch reference may be changed with the TCS button located on the pilot’s andcopilot’s control wheel, or by using the pitch wheel on the PC-400 AutopilotController.

Pitch attitude hold is canceled by selecting any vertical flight director mode orautomatic capture of a vertical mode.

Table 2-10-9. Pitch Attitude Hold Operating Limits


Autopilot Limit AfterEngagement

± 20°

TCS ± 20°

Pitch Wheel ± 20°



(3) Pitch Attitude Hold Mode Engage/Reset/Disengage Logic

(a) Engage Logic

Required valids:

• Flight Director

• On-side AHRU

• On-side MADC.

With the above conditions met, selecting a lateral flight director mode only,with no vertical flight director mode active, will place the aircraft in the pitchattitude hold mode of operation.

(b) Reset Logic


The pitch attitude hold mode will be automatically reset, if any of the followingconditions happen:

• Selecting any vertical flight director mode active



• Activation of the GC-550 CPL pushbutton.

(c) Disengage Logic


The pitch attitude hold mode will automatically disengage, if any of thefollowing conditions happen:







(4) Vertical Speed (VS) Hold Mode


The vertical speed hold mode is used to automatically maintain the aircraft at a pilotselected vertical speed reference. Mode activation will cancel all other verticalmodes except altitude preselect arm, and glideslope arm. Overspeed protectionbased on Vmo/Mmo speed limit is provided as a submode of vertical speed hold.Underspeed protection, in addition to AP disconnect at stick shaker activation isbeing considered at the time of the writing.

To initiate the mode, the pilot would maneuver the aircraft to the desired climb ordescent attitude, establish the vertical speed reference, and engage the mode.When the vertical speed mode is engaged, the following occurs:

• VS in green is annunciated on the PFD

• The vertical speed target is displayed above the vertical speed scale on the PFDand the vertical speed bug is displayed on the vertical speed scale.

The reference vertical speed is changed by turning the SPD knob on the GC-550Guidance Control Unit. If the autopilot is engaged after the mode is selected on, thereference vertical speed must be re-synchronized. The pilot can also change thevertical speed reference as a function of pressing and holding the TCS button andmanually flying the aircraft to a new vertical speed reference.

Input data for the vertical speed control law includes vertical speed target, actualaircraft vertical speed and normal acceleration.



180

160

220

240

7500

8500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

200 80 00

169 10000HDG VS

E29.92 IN

090 CRS

VOR2

15.0

HDG360

VOR2

NM

KTSGSPD

225

YDM

1000

20

80

1000

AP

ASEL

AD-50994@

Figure 2-10-27. Vertical Speed (VS) Hold Mode

Table 2-10-10. Vertical Speed Hold Operating Limits


VS Hold VS Engage Range 0 to ±6,000 ft/min

VS Hold Engage Error ± 30 ft/min

Pitch Limit ± 20°

Pitch Rate Limit f(TAS) 0.3 g’s max



(5) Vertical Speed (VS) Hold Mode Engage/Reset/Disengage Logic

(a) Engage Logic

Required valids:

• Flight Director

• On-side AHRU

• On-side MADC.

With the above conditions met, pressing the VS button on the GC-550Guidance Control Unit will engage the VS hold mode.

(b) Reset Logic


The VS mode will automatically be reset, if any of the following conditionshappen:


• Pressing the VS button on the GC-550 Guidance Control Unit

• Selecting any other vertical mode active



(c) Disengage Logic


The VS hold mode will automatically disengage, if any of the followingconditions happen:




• SG reversionary switching.



(6) Speed (SPD) Select Mode


The Speed Select mode is used to maintain a pilot selected speed reference. TheSPD control law is designed to fly to a selected altitude at a selected speed and toprovide limited overspeed/underspeed protection during climbs and descents.

The SPD mode is engaged by pressing the SPD button on the GC-550 GuidanceControl Unit. The desired speed reference is set by turning the SPD set knob on theGC-550 Guidance Control Unit. If the mode is engaged below 25,000 feet, an IASspeed reference is selected. If the aircraft climbs through the 25,100 feet, the IASreference is maintained. The pilot can also change the SPD reference by pressingand holding the TCS button and manually flying the aircraft to a new SPD reference.

If the mode is engaged above the 25,100 feet, a Mach reference is selected. If theaircraft descends through the 25,000 feet, the Mach reference is maintained. Thepilot may toggle between IAS/Mach at any time by pressing in on the SPD set knob.

When the SPD mode is engaged, the following occurs:

• IAS in green is annunciated on the PFD

• The speed target is displayed at the top of the airspeed scale on the PFD andthe speed target bug appears on the airspeed scale.

The SPD control law is designed to meet the following requirements:

• Not to exceed 0.3 g normal acceleration

• Not to exceed Vmo/Mmo

• Not to fly away from the selected airspeed reference

• Not to exceed 0.1 g longitudinal acceleration.



180

160

220

2408500

7500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

200 80 00

200 10000HDG IAS

E29.85 IN

090 CRS

VOR1

12.5

25

HDG001

VOR2

NM

KTSGSPD

225

YD

1500

20

80

-2000

AP

ASEL

AD-50995@

Figure 2-10-28. Speed Hold Mode



Table 2-10-11. Speed (SPD) Hold Mode Operating Limits


SPD Hold IAS Engage Range 120 knots to Vmo

IAS Hold Engage Error ± 5 knots

Pitch Limit ± 20°


Mach Engage Range .40 Mach

Mach Hold EngageError

± 0.01 Mach

Pitch Limit ± 20°


(7) Speed (SPD) Select Mode Engage/Reset/Disengage Logic

(a) Engage Logic

Required valids:

• Flight Director

• On-side AHRU

• On-side MADC.

With the above conditions met, pressing the SPD button on the GC-550Guidance Control Unit will engage the SPD Select mode.

(b) Reset Logic


The SPD mode will automatically be reset, if any of the following conditionshappen:


• Pressing the SPD button on the GC-550 Guidance Control Unit

• Selecting any other vertical mode active except altitude preselect arm orglideslope arm





(c) Disengage Logic


The SPD Select mode will automatically disengage, if any of the followingconditions happen:





(8) Flight Level Change (FLC, FLCH) Mode


The flight level climb or descent change mode is used to fly to a new altitudereference using predefined profiles. The flight level change control law is designedto provide limited overspeed/underspeed protection during climbs and descents. Thecontrol law is designed to meet the following requirements:

• Not to exceed 0.1 g normal acceleration

• Not to exceed Vmo/Mmo

• Not to fly away from the selected altitude

• Not to fly away from the profile speed

• Not to exceed 0.1 g longitudinal acceleration.

Flight level change provides an error signal from the reference speed and allows thesystem to maintain airspeed, Mach, or vertical speed and also track changes in theprofile speed target value.

The flight level change mode is selected by pressing the FLC button on the GC-550Guidance Control Unit.

Automatic selection of a climb or descent profile will be made by observing theaircraft altitude relative to the preselected altitude. The pilot will not be able toselect a speed reference with the SPD knob on the GC-550 Guidance Control Unit.Selection of this mode will cancel all other vertical flight director modes exceptaltitude preselect arm, or glideslope arm.

With FLC active, pressing and holding TCS will allow the pilot to maneuver theaircraft as he desires. When TCS is released, the aircraft will once again acquirethe climb/descent profile.

Input data used by the control law includes profile airspeed or Mach target, verticalspeed, actual airspeed or Mach, and aircraft accelerations.



The flight level change mode is canceled/and or inhibited by any of the followingconditions:

• FLC push button (push off)

• Selection of ALT, VS, or SPD modes

• Altitude preselect capture

• Glideslope capture

• Select Go-Around



220

200

260

28010500

9500

AD-51643@

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

N

S

33

15

3012

WE

246

21

3

240 100 00

240 37000HDG CLB

0.00 IN200 RA

316 CRS

VOR1

6.4

HDG360

ADF2

VOR1

NM

KTSGSPD

200

YD

DG1

800

20

80

-2000

AP

Figure 2-10-29. FLC Hold Mode



Table 2-10-12. Flight Level Change (FLC) Hold Mode Operating Limits

Mode Parameter Baro CorrectedAltitude (feet)

Value

FLC Hold Climb Speed 0-10,00010,000-12,00012,000-17,37717,377-37,000

240 knots240-270 knots270 knots0.56 Mach

Descent 0-37,000 -2,000 feet perminute(currently onchange. Newvalue TBD)

(9) Flight Level Change (FLC) Mode Engage/Reset/Disengage Logic

(a) Engage Logic

Required valids:

• Flight Director

• On-side AHRU

• On-side MADC.

With the above conditions met, pressing the FLC button on the GC-550Guidance Control Unit will engage the FLC mode.

(b) Reset Logic


The FLC mode will automatically be reset, if any of the following conditionshappen:


• Pressing the FLC button on the GC-550 Guidance Control Unit to turn offthe mode

• Selecting any other vertical mode active except altitude preselect arm orglideslope arm





(c) Disengage Logic


The FLC mode will automatically disengage, if any of the following conditionshappen:





(10) Altitude Preselect (ASEL) Mode


The ASEL mode is used in conjunction with another vertical mode to climb ordescend to a preselected altitude, automatically level off and maintain the barometricaltitude reference. Using the ASEL knob on the GC-550 Guidance Control Unit, thedesired barometric altitude is entered in the altitude alerter display window above thealtitude scale on the PFD.

The ASEL mode will arm automatically when the aircraft climbs or descends towardsthe desired altitude, when the following conditions are met:

• ASEL altitude is more than 250 feet from current altitude



• Aircraft is moving toward the target altitude

• Glideslope not captured.

The ASEL mode is canceled in altitude (ALT) hold or after glideslope capture.

The ASEL ARM mode is annunciated on the PFD. VS, SPD, FLC, or pitch hold canbe used to fly to the preselected altitude. When reaching the preselect bracketaltitude, the system automatically switches to the ALT SEL CAP mode and theprevious active pitch mode is canceled. A command is then generated toasymptotically capture the selected altitude.



The bracket altitude is defined as follows:

• When climbing towards the target altitude, the ASEL error is less than 2000 feetand computed vertical speed is less than actual vertical speed

• When descending towards the target altitude, the ASEL error is less than 3600feet and computed vertical speed is less than actual vertical speed.

The flare command generated during the altitude capture phase is a referenced VScommand generated by a linearized 0.06G approximation of the ASEL errorcomputation.

ASEL capture is annunciated on the PFD by a green ASEL at the vertical captureannunciation location. To indicate the transition to capture, ASEL will be enclosed ina white box for 5 seconds.

The aircraft will remain in the ASEL capture mode until both of the followingconditions exists:

• ALT error is less than 25 feet

• VS is less than 300 FPM.

At this time, the ASEL mode is dropped and the system is automatically placed inthe altitude hold mode.

The ASEL CAP mode will be dropped and ASEL ARM will be automaticallyreselected following ASEL knob motion.

Figure 2-10-30 illustrates a descent from 18,000 feet using the ALT SEL mode.



SELECTED ALTITUDE 15,000 FEET

18,000 FEET PRESENT ALTITUDE1

16,000 FEET2

FLARE3

4

AD-50996@

Figure 2-10-30. Altitude Preselect Mode Pictorial

The sequence of events described below are keyed to Figures 2-10-31 through2-10-34, to illustrate how the aircraft is flown to a preselected altitude, using theASEL mode.

• Pilot/copilot sets the selected altitude with the ASEL knob on the GC-550Guidance Control Unit. The digital readout of the ASEL altitude is displayed inthe ASEL window above the altitude scale on the PFD. (See Figure 2-10-31.)

• Engage the pitch hold (PIT), speed hold (IAS or MACH), flight level change(FLC) or pitch hold modes to descend toward the selected altitude. Altitudepreselect is automatically armed and annunciated. (See Figure 2-10-32.)

• The altitude flare point (ASEL CAP) is dependent on vertical speed. (See Figure2-10-33.)

• ASEL capture is dropped and ALT HOLD is automatically engaged.(See Figure 2-10-34).



180

160

220

24018500

17500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

200 180 00

169 15000HDG ALT

E29.92 IN

360 CRS

VOR1

44.5

HDG001

ADF2

VOR1

NM

KTSGSPD

225

YD

.390 M

M

20

80

AP

AD-50997@

Figure 2-10-31. Prior to Descent - Altitude Hold Mode



180

160

220

24016500

15500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

200 160 00

169 15000HDG VS

E29.92 IN

359 CRS

VOR1

42.5

25

HDG001

ADF2

VOR1

NM

KTSGSPD

225

YD

.390 M

M

-1500

20

80

-1500

AP

ASEL

AD-50999@

Figure 2-10-32. During Descent - ASEL Armed Mode



220

200

260

28015500

15000

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

240 153 20

169 15000HDG ASEL

29.92 IN

359 CRS

VOR1

40.5

HDG001

ADF2

VOR1

NM

MINTTG

9.5

YD

.470 M

M

-1000

40

00

-2000

AP

AD-50998@

Figure 2-10-33. Start of Flare - ASEL Capture



220

200

260

28015500

14500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

240 150 00

169 15000HDG ALT

29.92 IN

360 CRS

VOR1

38.5

HDG001

ADF2

VOR1

NM

MINTTG

9.0

YD

.470 M

M

20

80

-2000

AP

AD-51001@

Figure 2-10-34. Level at New Altitude - Altitude Hold Mode

The ASEL capture mode is canceled by:

• Moving the ASEL set knob on the GC-550 Guidance Control Unit

• Any other vertical mode selected on, or captured (except glideslope arm)

• Selecting go-around.



Table 2-10-13. Altitude Preselect (ASEL) Mode Operating Limits


ALT SEL ALT SEL CaptureRange

0 to 50,000 ft

ALT Capture Error ± 25 ft

Pitch Limit ± 20°


(11) Altitude Preselect (ASEL) Mode Engage/Reset/Disengage Logic

(a) Engage Logic

Required valids:

• Flight Director

• On-side AHRU

• On-side MADC.

With the above conditions met, and all of the following conditions met, theASEL mode will automatically arm:

• Decreasing altitude error


• Not altitude hold

• Glideslope not capture or track


• The target altitude is at least 250 feet from present altitude.

(b) Reset Logic


The ASEL mode will automatically be reset, if any of the following conditionshappen:

• Altitude select capture


• Selecting any other vertical mode active except glideslope arm





(c) Disengage Logic


The ASEL mode will automatically disengage, if any of the following conditionshappen:





(12) Altitude Hold (ALT) Mode


The altitude hold mode is a vertical axis flight director mode used to maintain abarometric altitude reference. The vertical command for altitude hold is displayed onthe flight director pitch command bar on the PFD. To fly utilizing altitude hold, thepilot would:

• Establish the aircraft in straight and level flight

• Press the ALT button on the GC-550 Guidance Control Unit.

At this time, the green ALT annunciator is displayed on the PFD while altitude hold isactive. The vertical axis of the flight director will maintain the barometric altitude atthe time of mode engagement. The reference altitude may be changed by usingTCS to maneuver to a new altitude and then releasing the TCS button. Selecting theALT mode on will cancel any other previously selected vertical mode.

The pilot can change the altitude hold reference by pressing and holding the TCSbutton and manually fly the aircraft to a new altitude reference.

Altitude Hold can also be entered automatically, as a function of ASEL capture, flareand level off.

The ALT hold mode is canceled by:

• Pressing the ALT button on the GC-550 Guidance Control Unit

• Selecting any other vertical mode on the GC-550 Guidance Control Unit



• Pitch wheel movement in either direction on the PC-400 Autopilot Controller.



180

160

220

24010500

9500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

200 100 00

169 10000LNAV ALT

E30.01 IN

246 DTK

FMS

13.1

25

HDG245

VOR1

NM

KTSGSPD

215

YDM

20

80

-2000

AP

AD-51000@

Figure 2-10-35. Altitude Hold Mode

Table 2-10-14. Altitude Hold (ALT) Mode Operating Limits


ALT ALT Capture Range 0 to 50,000 ft

ALT Hold CaptureError

± 20 ft

Pitch Limit ± 20°




(13) Altitude Hold (ALT) Mode Engage/Reset/Disengage Logic

(a) Engage Logic

Required valids:

• Flight Director

• On-side AHRU

• On-side MADC.

With the above conditions met, pressing the ALT button on the GC-550Guidance Control Unit will engage the altitude hold mode.

(b) Reset Logic


The ALT mode will automatically be reset, if any of the following conditionshappen:

• Altitude select arm





• Pitch wheel movement on the PC-400 Autopilot Controller.

(c) Disengage Logic


The ALT mode will automatically disengage, if any of the following conditionshappen:







(14) ILS Approach (APR) Mode

See Figure 2-10-36 through 2-10-40 and Table 2-10-15.

The glideslope portion of the approach mode is used for the automatic intercept,capture and tracking of the glideslope beam. The beam is used to guide the aircraftdown to the runway in a linear descent. Typical glideslope beam angles varybetween two and three degrees, dependent on local terrain. When the glideslopemode is used as the vertical portion of the localizer approach mode, it allows thepilot to fly a fully coupled ILS approach. The mode is interlocked, so that glideslopecapture is inhibited until localizer capture has occurred.

With the localizer captured, and outside the normal glideslope capture limits, thePFD will annunciate the following modes:

• LOC in green

• GS in white

• Any other vertical mode in use at this time will also be displayed.

As the aircraft approaches the glideslope beam, the vertical beam sensor (VBS)monitors TAS, vertical speed, and glideslope deviation in determining the correctcapture point. At glideslope capture, the computer drops any other vertical modethat was in use, and automatically generates a pitch command to smoothly track theglideslope beam.

At this time, the PFD will annunciate LOC in green and GS in green (thisannunciation will be enclosed with a white box for 5 seconds).

The NAV and APR mode selector buttons on the GC-550 Guidance Control Unit areannunciated.



090°INBOUNDCOURSE

LOCALIZERTX

AD-50988@

RUNWAY



Figure 2-10-36. ILS Approach Arm Pictorial



140

120

180

2003500

2500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

160 30 00

169 3000HDG ALT

E

29.96 IN200 RA

090 CRS

ILS1

8.0

HDG045

ADF1

NM

KTSGSPD

150

YDM

1560

20

80

-2000

AP

GSLOC

R

AD-50990@

Figure 2-10-37. ILS Approach Arm

Gain programming is performed on the glideslope signal to compensate for theaircraft closing on the glideslope antenna, and beam convergence caused by thedirectional properties of the glideslope antenna. Glideslope programming is normallyaccomplished as a function of the change in radio altitude on the approach.

If radio altitude is not valid, or not available, glideslope gain programming starts atglideslope capture and will run down as a function of TAS and time. At the middlemarker, the gain is synchronized to a preset value for the remainder of the approach.

Input data used in the glideslope control law includes glideslope deviation, radioaltitude, airspeed, vertical speed and middle marker.



The approach mode is canceled by:

• Pressing the APR or NAV buttons on the GC-550 Guidance Control Unit

• Loss of ILS localizer or glideslope data


• Selecting HDG on the GC-550 Guidance Control Unit

• Selecting any other vertical mode active.

AD-50982@

RUNWAY

GLIDESLOPE BEAM

TYPICAL CAPTURE POINT IS1/3 DOT ON GS SCALE

Figure 2-10-38. ILS Approach Capture Pictorial



140

120

180

2003500

250030

30

30

30

3

2

1

0

1

2

3

160 30 00

169 3000LOC ALT

E

29.96 IN200 RA

090 CRS

ILS1

6.8

HDG045

ADF1

NM

KTSGSPD

140

YDM

1500

20

80

-2000

AP

GS

R

AP

AD-50991@

Figure 2-10-39. ILS Approach (LOC) Capture



120

100

140

160

2200

1500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

130 17 60

169 3000LOC GS

AP

29.92 IN200 RA

090 CRS

ILS1

1.3

HDG045

ADF1

NM

KTSGSPD

100

YDM

-600

200

40

80

-2000

AP

R

M

AD-50993@

Figure 2-10-40. ILS Approach (APR) Mode Track

Table 2-10-15. ILS Approach (APR) Mode Operating Limits


APR Glideslope Capture < 150 mV glideslope deviation

Capture Point Function of GS deviation, verticalspeed, and TAS

Pitch Command Limit 10°


Gain Programming Function of radio altitude, TAS,and GS deviation



(15) ILS Approach (APR) Mode Engage/Reset/Disengage Logic

(a) APR Arm Engage Logic

Required valids:

• Flight Director

• On-side AHRU

• On-side NAV source.

With the above conditions met, and the NAV source tuned to an ILSfrequency, pressing the APR button on the GC-550 Guidance Control Unit willarm the ILS approach mode.

(b) APR Arm Reset Logic


APR arm will automatically be reset, if any of the following conditions happen:


• Pressing the NAV or HDG button on the GC-550 Guidance Control Unit





(c) Disengage Logic


The APR arm mode will automatically disengage, if any of the followingconditions happen:



• Not tuned to an ILS frequency




(d) ILS Approach Capture Logic

Required valids:

• Flight Director system

• On-side AHRU


With the above conditions met, glideslope armed and the VBS has tripped, theILS Approach mode will automatically transition from arm to capture.

(e) ILS Capture Reset Logic

ILS capture will automatically be reset, if any of the following conditionshappen:


• Pressing the NAV or HDG button on the GC-550 Guidance Control Unit





(f) ILS Capture Disengage Logic

The ILS capture mode will automatically disengage, if any of the followingconditions happen:


• Not tuned to an ILS frequency


• NAV source not valid for 5 seconds


(g) ILS Track Engage Logic

Required valids:

• Flight Director

• On-side AHRU


With the above conditions met, glideslope is captured and localizer is trackand radio altitude is less than 1200 ft, the ILS mode will automaticallytransition from capture to track.



ILS track will automatically be reset for the same conditions as ILS CaptureReset.

The ILS track mode will automatically disengage for the same reasons as ILScapture disengage.

NOTE: If glideslope deviation becomes invalid, the commandcue will bias out of view but the mode annunciatorstays on. If the glideslope deviation remains invalidfor more than 5 seconds, the mode will be canceled.

(16) Go-Around (GA) Mode (Wings Level)

See Figure 2-10-41.

The Go-Around mode will be selected by pressing the remote mounted go aroundbutton. The IC-600 will provide flight director bar and autopilot pitch commands toallow Speed-on-Elevator aircraft control.

The Coupled Go Around mode will initially command a minimum speed of 1.23 Vskts to arrest the aircraft descent and begin the climb out. The reference speed willbe latched on mode entry. The mode will ensure that a minimum of altitude is lostgiven that the aircraft has sufficient energy to maintain speed and altitude. In theevent of insufficient energy, the aircraft will maintain altitude until 1.23 Vs kts isreached, at which point the mode will maintain speed. The turn around will be doneutilizing energy management to command a minimum of 75% of the achievableacceleration or the acceleration needed to meet the 1.23 Vs kts speed target. Themode will achieve a pitch up attitude that is limited by a comfort limit of 12 degrees,at which point the aircraft will be allowed to accelerate depending on the energyavailable.



120

160

180

1500

2500

2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

N

S

33

15

3012

WE

246

21

3

8

14 21 40

151 3800ROL GA

29.92 IN200 RA

360 CRS

ILS1

5.2

HDG360

NM

KTSGSPD

148

YD

700

190

20

60

-2000

AP

MIN

AD-51642@

GS

ASEL

ADF2

VOR1

6

Figure 2-10-41. Go-Around Mode (Wings Level)

The Go-Around mode can be canceled or inhibited by any of the following:

• MADC source selection change

• Selection of any other vertical flight director mode

• ASEL capture

• Pressing TCS

• Radio altitude valid and radio altitude > 2500 ft

• Radio altitude invalid and baro corrected altitude > 1500 ft.



(17) Go-Around (GA) Mode Engage/Reset/Disengage Logic

(a) Engage Logic

Required valids:

• Flight Director

• On-side MADC

• On-side AHRU.

With the above conditions met, pressing the remote GA button on the pilot’s orcopilot’s throttle will engage the GA mode.

(b) Reset Logic


The GA mode will automatically be reset, if any of the following conditionshappen:


• Pushing TCS



• MADC source selection change

• Transition to capture phase of altitude preselect mode (provided that theselected altitude is greater than 400 ft above coupled go around modeengagement altitude.

(c) Disengage Logic


The GA mode will automatically disengage, if any of the following conditionshappen:






(18) Windshear Mode (WSHR)

The windshear (WSHR) system utilizes a combination of detection and guidance.The windshear detection computer will annunciate an amber or red WDSHEAR onthe PFD for increasing performance or decreasing performance windshearsrespectively. See Figure 2-10-41.

Anytime an amber or red WDSHEAR is annunciated, the pilot is expected to give fullthrottle, activate the WSHR mode and execute a missed approach while following theWSHR guidance.

The WSHR mode is activated manually by pressing the GA button while an amber orred WDSHEAR annunciation exists. It is automatically activated from the TO or GAmodes if an amber or red WDSHEAR occurs. It is also automatically activated fromany mode if a red WDSHEAR occurs and the throttles are greater than 80% of thethrottle lever angular travel.

When the windshear mode is activated, the autopilot is disconnected and thecommand bars will be displayed. The command bars will guide the aircraft gainenergy or conserve energy depending on the conditions of the airmass the aircraft isin. This energy management is done first by maintaining a slightly negative (climb)flight path angle. This will continue until the aircraft has either reached excessenergy and begins to control to a speed target for climb out or until the aircraft haspitched up near stickshaker while trying to maintain the flight path.

If the aircraft has reached excess energy, the aircraft will switch from flight pathcontrol to speed control and allow the aircraft to reach whatever climb rate it canwhile maintaining airspeed. Excess energy is defined as the latched speed at timeof WSHR engagement which is the greater of 1.3Vs + 30kts or current IAS. Thisspeed latch is increased by 20 knots if an increasing performance (amber) windshearis last detected.

If the aircraft has pitched up while trying to maintain flight path angle, the WSHRguidance will control to stickshaker until a climb can again be achieved.



80

100

160

2000

1500

140 2020

1010

1010

2020

3030

3030

3

2

1

0

1

2

3

120 18 20

169 3800ROL WSHR

AP

29.29 IN200 RA

360 CRS

ILS1

6.7

HDG360

NM

KTSGSPD

122

1830

40

00

-2000

WDSHEAR

GS

8

ADF2

VOR1

AD-51727@

Figure 2-10-42. Windshear Mode

(19) Windshear (WSHR) Mode Engage/Reset/Disengage Logic

(a) Engage Logic

Required valids:

• Caution or warning signal from the GPWS Computer

• Stall Warning Computer

• On-side MADC

• On-side AHRU

• Radio altitude and below 1500 ft.

With the above conditions met, pressing the remote GA button on the pilot’s orcopilot’s throttle will engage the windshear mode.



The windshear mode will automatically be entered when the followingconditions happen:

• In GA or TO mode and a valid state exists for an increasing performanceshear (windshear caution) or a decreasing performance shear (windshearwarning)

• Throttle lever position exceeds 80% and a valid state exists specifically fora decreasing performance shear (windshear warning)

(b) Reset Logic


The windshear mode will automatically be reset, if any of the followingconditions happen:

• Selection of FLC, VS, SPD and ALT modes

• The go around button is pressed without a windshear caution or warning

• MADC source selection change without a windshear or caution warning

• AHRS source selection change without a windshear or caution warning



(c) Disengage Logic


The windshear mode will automatically disengage, if any of the followingconditions happen:

• Stall Warning Computer not valid



• Radio altitude is greater than 1500 ft

• No discrete from the windshear computer exists.





2.11 AUTOPILOT/YAW DAMPER SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-1

1. General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-1

A. Autopilot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-1

B. Yaw Damper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-2

C. AP/YD System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-2

2. Component Descriptions and Locations . . . . . . . . . . . . . . . . . . . . 2-11-4


B. GC-550 Guidance Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-7

(1) Autopilot Engage Pushbutton . . . . . . . . . . . . . . . . . . . . . . . 2-11-9

(2) Yaw Damper Engage Pushbutton . . . . . . . . . . . . . . . . . . . . 2-11-9

(3) CPL Pushbutton . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-9

C. PC-400 Autopilot Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-10

D. SM-200 Servo Drive and SB-201 Servo Bracket . . . . . . . . . . . . . . 2-11-12

3. Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-15

A. Modes of Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-15

(1) Heading Hold and Wings Level . . . . . . . . . . . . . . . . . . . . . 2-11-15

(2) Roll Hold . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-15


(4) Flight Director Couple and Lift Compensation . . . . . . . . . 2-11-16

(5) Turn Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-17

(6) Pitch Wheel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-17

(7) Touch Control Steering (TCS) . . . . . . . . . . . . . . . . . . . . . . 2-11-17

B. Autopilot/Yaw Damper Engage Logic . . . . . . . . . . . . . . . . . . . . . . 2-11-23

(1) Yaw Damper Engagement . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-23

(2) Autopilot Engagement . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-24

C. Roll Axis Autopilot Servo Loop . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-27

(1) SM-200 Roll Servo Drive and Bracket . . . . . . . . . . . . . . . . 2-11-27


D. Pitch Axis Autopilot Servo Loop . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-33

(1) SM-200 Elevator Servo Drive and Bracket . . . . . . . . . . . . . 2-11-33


E. Pitch Axis Autopilot Trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-36

F. Yaw Damper Rudder Axis Servo Loop . . . . . . . . . . . . . . . . . . . . . 2-11-39


(2) SM-200 Servo Drive and Bracket . . . . . . . . . . . . . . . . . . . 2-11-40

4. Fault Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-43

A. Autopilot/Yaw Damper Monitoring Overview . . . . . . . . . . . . . . . . 2-11-43





B. Hardover Malfunction Protection . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-43

C. System Response to Failures . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-43

D. Monitor Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-44

(1) Pitch Servo Position Monitor . . . . . . . . . . . . . . . . . . . . . . 2-11-44

(2) Primary Pitch Attitude Comparison . . . . . . . . . . . . . . . . . 2-11-44

(3) Secondary Pitch Attitude Comparision Monitor . . . . . . . . 2-11-44

(4) Normal Accelertion Monitor . . . . . . . . . . . . . . . . . . . . . . . 2-11-44

(5) Roll Servo Position Monitor . . . . . . . . . . . . . . . . . . . . . . . 2-11-44

(6) Primary Roll Attitude Comparison Monitor . . . . . . . . . . . . 2-11-45

(7) Secondary Roll Attitude Comparison Monitor . . . . . . . . . . 2-11-45

(8) Roll Rate Monitor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-45

(9) Yaw Servo Position Monitor . . . . . . . . . . . . . . . . . . . . . . . 2-11-45

(10) Auto Trim Runaway Monitor . . . . . . . . . . . . . . . . . . . . . . . 2-11-45

(11) Auto Trim Inoperative Monitor . . . . . . . . . . . . . . . . . . . . . 2-11-45

(12) Autopilot/Yaw Damper Disconnect Monitor . . . . . . . . . . . . 2-11-46




FIGURE/TITLE PAGE



Figure 2-11-2. GC-550 Guidance Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-7


Figure 2-11-4. SM-200 Servo Drive and SB-201 Servo Bracket . . . . . . . . . . . . . . . . . . . 2-11-12

Figure 2-11-5. Pilot’s Autopilot/Yaw Damper Interface . . . . . . . . . . . . . . . . . . . . . . . 2-11-19/20

Figure 2-11-6. Copilot’s Autopilot/Yaw Damper Interface . . . . . . . . . . . . . . . . . . . . . 2-11-21/22

Figure 2-11-7. Autopilot/Yaw Damper Engage Logic . . . . . . . . . . . . . . . . . . . . . . . . 2-11-25/26

Figure 2-11-8. Autopilot Roll Axis Servo Loop . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-31/32

Figure 2-11-9. Pitch Autopilot Servo Loop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-37/38

Figure 2-11-10. Yaw Damper Servo Loop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-41/42

List of Tables

TABLE/TITLE PAGE

Table 2-11-1. IC-600 Integrated Avionics Computer (Autopilot/Yaw Damper Function)

Leading Particulars . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-5

Table 2-11-2. GC-550 Flight Guidance Control Unit Leading Particulars . . . . . . . . . . . . 2-11-7


Table 2-11-4. SM-200 Servo Drive and SB-201 Servo Bracket Leading Particulars . . . . 2-11-13

Table 2-11-5. Autopilot Roll Axis Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-29

Table 2-11-6. Pitch Channel Axis Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-11-35



(Blank Page)



SECTION 2.11AUTOPILOT/YAW DAMPER SYSTEM

1. General

The PRIMUS® 1000 Display and Flight Guidance System features a single autopilot/yawdamper system that is designed around a distributed processor architecture which utilizesindependent hardware elements to perform the aircraft control and monitor functions. Themonitor function is performed in the primary (flight director/EDS) processor within the IC-600IAC No. 1, while the aircraft control function is performed in the secondary (autopilot/yawdamper) processor.

This architecture ensures that any single failure does not cause a condition which would preventcontinued safe flight and landing of the aircraft. In the event of a primary processor failure, thesecondary processor would become unavailable for use since the monitor functions are housedin the primary processor. If however, the secondary processor failed, the functions of theprimary processor would still be available.

The IC-600 IAC No. 1 Integrated Avionics Computer houses the autopilot/yaw damper. Only thepilots IC-600 is connected to the servos, consequently this configuration only provides for singleautopilot/yaw damper operation.

The PRIMUS® 1000 autopilot/yaw damper system requires that both AHRS in the aircraft beoperating and be valid. The autopilot/yaw damper is not designed for single AHRS operation.

The primary processor provides dedicated disconnect hardware for the monitor function. Thisallows either processor to force a disconnect of the autopilot and yaw damper. All automaticdisconnects, which result from monitor trips, will be stored in non-volatile memory for later recallduring ground maintenance test.

A. Autopilot

The PRIMUS® 1000 autopilot is housed in the pilot’s IC-600 Integrated Avionics Computer(IAC) and is of a fail passive design featuring digital attitude and servo loops. Theautopilot provides attitude stabilization and tracking of pitch and roll steering commandsfrom the flight director. The autopilot is not aware of which flight director mode(s), if any,are active. The autopilot simply tracks the pitch and roll steering commands as attitudechanges.

The autopilot provides aircraft stabilization around a pilot selected reference. With theautopilot engaged, short term transient disturbances are automatically corrected. As theaircraft is moved away from its reference by a disturbance, the autopilot will work to stopthe aircraft from moving away and return it to its reference position/attitude.



The pitch axis autopilot trim function resides in the pilots IC-600 IAC and works to maintainaircraft pitch attitude against long term attitude disturbances such as fuel burn andpassenger movement. Activation of the manual electric trim switches will cause theautopilot to disengage.

For the autopilot to do its job, it considers the following data:

• What is the pilots desired attitude reference?

• What is the aircrafts actual attitude?

• If there is a difference between desired and actual attitude, correct for the differenceand control the rate at which the correction takes place.

B. Yaw Damper

The yaw damper computes servo commands based on sensor input data only. It providesyaw rate damping and makes no effort to control the flight path of the aircraft. While theyaw damper can be engaged without the autopilot, the autopilot cannot be engaged withoutthe yaw damper.

Servo position reference is synchronized to zero at engagement and is constantly washedout to ensure that steady state rudder forces are zero. If the rudder trim position changesdue to pilot input or aircraft configuration changes the rudder washes out the steady stateforce and allows rudder servo re-synchronization.

C. AP/YD System

The PRIMUS® 1000 autopilot/yaw damper system is comprised of the following LineReplaceable Units (LRUs).

• IC-600 IAC (Pilot’s)

• GC-550 Guidance Control Panel

• SM-200 Servos and Servo Brackets (aileron, elevator, rudder)

• PC-400 Autopilot Controller

• Aircraft Primary Trim System.

The autopilot/yaw damper requires inputs from the following sensors:

• AHRS (Pilot’s and Copilot’s)

• AZ-850 Micro Air Data Computer.

Autopilot modes of operation are listed below:

• Heading Hold

• Roll Hold

• Pitch Attitude Hold

• Flight Director Coupled.



The AP/YD off discrete output logic provides a 3.0-second AP off output for the warninghorn for normal autopilot disconnects, and a continuous output for any automaticdisconnect. The continuous output can be reset by holding the AP disconnect on thecontrol wheel for more than 1 second.





Figure 2-11-1 shows a graphical view of the IC-600 Integrated Avionics Computer. TwoIC-600 Integrated Avionics Computers are located under the cockpit floor behind the crew.Table 2-11-1 provides items and specifications that are particular to the computer.

AD-33449@




Table 2-11-1. IC-600 Integrated Avionics Computer(Autopilot/Yaw Damper Function) Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . 7.62 in. (193.55 mm)

• Width . . . . . . . . . . . . . . . . . . 4.13 in. (104.90 mm)

• Length . . . . . . . . . . . . . . . . . 16.45 in. (418.83 mm)

Weight (maximum):

• With Autopilot . . . . . . . . . . . . 15.5 lb (7.05 kg)

• Without Autopilot . . . . . . . . . . 15.0 lb (6.82 kg)

Power Requirements (with Autopilot):

• Continuous . . . . . . . . . . . . . . 28 V dc, 50 W (max)

• In-Rush . . . . . . . . . . . . . . . . 28 V dc (0.5 sec), 200 W (max)

• Servo Power . . . . . . . . . . . . . 28 V dc, 210 W (max)/112 W (nom)

Power Requirements (without Autopilot):

• Continuous . . . . . . . . . . . . . . 28 V dc, 50 W (max)

• In-Rush . . . . . . . . . . . . . . . . 28 V dc (0.5 sec), 200 W (max)

Mating Connectors:

• J1, J2 . . . . . . . . . . . . . . . . . . ITT Cannon Part No.DPX2MA-A106P-A106P-33B-0001

NOTE: Sunbank backshell (4) requiredPart No. J1560-12-2

Mounting . . . . . . . . . . . . . . . . . . . Tray, HPN 7017095-903

The primary component of the PRIMUS® 1000 autopilot/yaw damper system is the pilot’sIC-600 Integrated Avionics Computer (IAC). The autopilot/yaw damper processor in thepilot’s IC-600 IAC provides pitch and roll attitude commands through control of elevator,aileron and trim servos, as well as yaw rate terms to the rudder servo. The autopilot tracksthe pitch and roll attitude commands from the flight director to provide computed flight pathsteering. The CPL switch on the GC-550 Guidance Control Unit allows the pilot to selectwhich flight director (pilot/copilot) is coupled to the autopilot.



In addition to providing stabilization around a pilot defined reference, the autopilotprocessor also provides:

• Power up/start up initialization

• Engage/disengage logic

• Airspeed gain computations

• I/O data management

• ARINC 429 communications

• Continuous testing functions.



B. GC-550 Guidance Control Unit

Figure 2-11-2 shows a graphical view of the GC-550 Guidance Control Unit. The GC-550Guidance Control Unit is located in the center of the glareshield. Table 2-11-2 providesitems and specifications that are particular to the controller.

CRS 1

PUSH SYNC PUSH SYNC

HDG SPD

PUSH IAS/M

ASEL CRS 2

PUSH SYNC

FD1 HDG

BNK

APR

NAV AP

CPL

YD

SPD FD2ALT

VS

FLC

AD-50630@

Figure 2-11-2. GC-550 Guidance Control Unit

Table 2-11-2. GC-550 Flight Guidance Control Unit Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.25 in. (57.15 mm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.65 in. (295.91 mm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.51 in. (114.51 mm)


Power Requirements:




Table 2-11-2. GC-550 Flight Guidance Control Unit Leading Particulars

Item Specification


• Knobs










• Setscrews



• Lamps



Mating Connector:

• J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS27473E20B-35S




The GC-550 Guidance Control Unit provides the means for the pilot to engage/disengagethe autopilot and yaw damper and select which flight director is commanding the autopilot.

(1) Autopilot Engage Pushbutton

Autopilot engagement is primarily controlled through this pushbutton. If all engagelogic is valid, pressing this switch will engage the autopilot and yaw damper. Withthe autopilot engaged, pressing this switch will disengage the autopilot only. Theyaw damper will remain engaged.

(2) Yaw Damper Engage Pushbutton

Yaw Damper engagement is primarily controlled through this pushbutton. If allengage logic is valid, pressing this pushbutton will engage the yaw damper only.With the autopilot engaged, pressing this pushbutton will disengage the yaw damperand autopilot.

(3) CPL Pushbutton

This pushbutton is used to select which flight director is commanding the autopilot.Only one flight director at a time can be selected. Transition of this switch will cancelall flight director modes independent of autopilot operation. After the flight directortransfer has taken place, flight director modes can be re-selected by the pilot.



C. PC-400 Autopilot Controller

The PC-400 Autopilot Controller provides the autopilot with TURN knob and PITCH wheelinputs. The controller is located on the pedestal with Dzus rails with four, one-quarter turnfasteners. Figure 2-11-3 shows a graphical view of the PC-400 Autopilot Controller.Leading particulars for the PC-400 Autopilot Controller are listed in Table 2-11-3.

TURNDESCEND

CLIMB

PITCH

AD-50485@

Figure 2-11-3. PC-400 Autopilot Controller



Table 2-11-3. PC-400 Autopilot Controller Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.625 in. (6.67 cm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.750 in. (14.60 cm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.150 in. (15.62 cm)


Power Requirements:

• Instrument Lighting . . . . . . . . . . . . . . . . . . . . . . . . . 0 to 5 V ac or dc

• Mode Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . +28 V dc


• Knob, Turn . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HPN 337136-1

• Setscrew, Bottom (Hex Socket, 8-32 x 5/8", cuppoint) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

HPN 0455-284

• Setscrew, Side (Hex Socket, 8-32 x 3/16", cup point) HPN 0455-274

• Lamp, Clear (Type 7341) . . . . . . . . . . . . . . . . . . . . HPN 0635-22

Mating Connector:

• J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS3116F20-41S

The pilot may input roll commands with the TURN knob or change pitch attitude referencewith the PITCH wheel with the autopilot engaged.

The TURN knob drives a center tap potentiometer and cam mechanism. The cammechanism provides a no command mechanical detent and command sensing (out ofdetent) signal to the IC-600 IAC. The PITCH wheel drives a tachometer generator whichprovides a CLIMB or DESCEND dc voltage output proportional to the rotation of the wheel.



D. SM-200 Servo Drive and SB-201 Servo Bracket

Figure 2-11-4 shows a graphical view of the SB-200 Servo Drive and SB-201 ServoBracket. The aileron and elevator servos are located under the cockpit floor in betweenthe pilot and copilots seats. The rudder servo is located under the cockpit floor just behindthe engines. Table 2-11-4 provides items and specifications that are particular to the driveand bracket.

AD-34765@

Figure 2-11-4. SM-200 Servo Drive and SB-201 Servo Bracket



Table 2-11-4. SM-200 Servo Drive and SB-201 Servo Bracket Leading Particulars

Item Specification


• Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.970 in. (10.08 cm)

• Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.065 in. (12.88 cm)

• Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.825 in. (22.43 cm)


Power Requirement:

• Clutch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 V dc, 19 W (max)

• Motor Stall . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 V dc, 56 W

• Synchro . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26 V ac, 400 Hz, 1 VA

Small Torque . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Up to 160 lb-in.

Resistance Values:

• Clutch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57.5 ± 6 ohms

• Tachometer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 140 ohms (max)

• Motor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20.7 ± 2.6 ohms

NOTE: Motor resistance value is calculated by accurately measuringapplied voltage and load current. When attempting to measurethis resistance with an ohmmeter, the value may vary between18 and several hundred ohms, depending upon brush position andthe quantity of brush dust.


• Drive to Bracket Retaining Screws (4) . . . . . . . . . . HPN 4011086

• Cable Keepers (4) . . . . . . . . . . . . . . . . . . . . . . . . HPN 2518330

• Retaining Plate (1) . . . . . . . . . . . . . . . . . . . . . . . . HPN 2518332

• Screw, Cable Capture . . . . . . . . . . . . . . . . . . . . . . HPN 2554911-1

• Screw, Plate Retaining . . . . . . . . . . . . . . . . . . . . . HPN 0457-242

Mating Connector:

• J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PT06E-14-19S(SR)

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hard Mount



The SM-200 Servo Drive translates electrical inputs into clutched rotational mechanicaloutputs. Tachometer rate signals are fed back to the IC-600 IAC servo amplifier to null thecommand signal. Leading particulars are supplied in Table 2-11-4.

The SB-201 Bracket is firmly bolted to the aircraft airframe and the drum is connected tothe aircraft primary control rigging through cables. The SM-200 Servo Drive, with a splineoutput on the clutch, mates with the drum and bracket and may be removed from the drumand bracket without disturbing the aircraft rigging.



3. Operation

The PRIMUS® 1000 autopilot/yaw damper interface is shown in Figures 2-11-5 and 2-11-6.Figure 2-11-5 is the pilot’s side interface, and Figure 2-11-6 is the copilot’s side interface. Thisshows major signal flow between autopilot/yaw damper LRUs. For complete wiring interfacedata, refer to SECTION 3 of this manual.

A. Modes of Operation

The PRIMUS® 1000 autopilot has five modes of operation. These modes are HeadingHold and Wings Level, Roll Hold, Pitch Attitude Hold, Flight Director Coupled and TurnKnob.

(1) Heading Hold and Wings Level

Heading Hold is defined as the basic lateral default autopilot mode which isannuniciated as ROL on the PFD. It is defined as follows:

• Autopilot engaged

• Bank angle less than 6°

• No lateral flight director mode active.

If the conditions listed above are satisfied, the autopilot will roll the aircraft to awings level attitude. When the aircraft’s roll attitude is less than 3° plus 3 seconds,the heading hold mode is automatically engaged.

(2) Roll Hold

The roll hold mode is recognized as being active when the following conditions aresatisfied:

• Autopilot engaged

• No lateral flight director mode active

• The aircraft’s bank angle is greater than 6° but less than 35°.

Touch Control Steering (TCS) may be used to initiate a roll maneuver, while theautopilot is engaged. When all of the above conditions are satisfied, the autopilotwill maintain the prescribed bank angle. If TCS is released at bank angles greaterthan 35°, the autopilot will roll the aircraft to 35° and maintain the bank angle.

If TCS is released at bank angles less than 6°, the autopilot will revert to a wingslevel condition and then heading hold which annunciates as ROL on the PFD.



(3) Pitch Attitude Hold

Pitch attitude hold is the basic vertical mode of the autopilot whichs annunciates asPIT on the PFD. It automatically becomes active when:

• The autopilot is engaged or

• A lateral flight director mode is active and no vertical flight director mode isactive.

The position of the pitch command bar on the PFD provides the pilot with areference of aircraft pitch attitude at the moment the autopilot is engaged. This pitchattitude reference can be changed as a function of Touch Control Steering (TCS) oruse of the pitch wheel.

While in pitch attitude hold, pressing and holding the TCS switch on the controlcolumn will disengage the elevator and aileron servo clutches and synchronize theautopilot pitch reference to existing aircraft pitch attitude. The pilot can nowmanually fly the aircraft to a new pitch attitude reference and the autopilot memorywill synchronize to it. Releasing the TCS button will re-engage the elevator andaileron servo clutches and the pitch axis of the autopilot will provide stabilizationaround this new reference.

(4) Flight Director Couple and Lift Compensation

With the autopilot engaged, anytime a flight director mode is selected on, thecomputed steering command (attitude change) is transmitted to the autopilot. Theautopilot in turn develops a servo loop command to drive the appropriate flightcontrol surface to satisfy the flight director input. This coupling of flight director andautopilot allows hands off automatic flight path steering throughout the aircraft’s flightregime.

Just as pilots are taught to keep the nose of the aircraft up when making a turn, theautopilot must have the same ability for the same reason. When banking an aircraftto make a turn, lift is lost on one wing. This loss of lift results in the aircraft losingaltitude. To compensate for this manually, the pilot would apply a slight backpressure to the control column to hold the nose up and not lose altitude in the turn.

The autopilot will accomplish this automatically through a design feature referred toas "lift compensation". This is done creating a term that is equal to the cosine of thebank angle subtracted from 1.0 and applying this term to the pitch axis of theautopilot. This in effect keeps the nose of the aircraft in the proper attitude to notlose altitude as the turn is made.



(5) Turn Knob

Rotation of the TURN knob out of detent results in a roll command whichannunciates ROL on the PFD. The roll angle is proportional to and in the directionof the TURN knob rotation. The TURN knob controls a detent switch andpotentiometer to provide roll command to the IC-600 IAC. The TURN knob must bein detent (center position) before the autopilot can be engaged. Rotation of theTURN knob out of detent cancels any lateral mode selected.

(6) Pitch Wheel

Rotation of the pitch wheel which annunciates PIT on the PFD results in a change ofpitch attitude proportional to the rotation of the wheel and in the direction of wheelmovement. The pitch wheel provides rate limited pitch commands in pitch holdmode. The pitch thumb wheel provides a tachometer output which is applied to thePilot’s IC-600 IAC. Pitch wheel operation is inhibited in GS. Pitch wheel operationwill cancel all other modes.

(7) Touch Control Steering (TCS)

TCS allows the pilot to momentarily disengage the autopilot and manually fly theaircraft to a new pitch/roll attitude reference.

The TCS switches are located on the pilot’s and copilot’s control wheels. Wheneither switch is activated and held, the following will occur:

• Autopilot clutches (aileron and elevator) disengage

• Autopilot pitch axis memory will synchronize to current aircraft position. When thepilot completes the maneuver and releases the TCS switch, the autopilot clutchre-engages and the autopilot now holds the new pitch attitude reference.Depending upon the bank angle at TCS release, the autopilot will go into eitherwings level or roll hold mode.



(Blank Page)



8

9

AZ-850 MADCNO. 1

ARINC 429BUS NO.1

ARINC 429BUS NO.2

(H)(L)

(H)(L)

6364

TO HORIZONTALSTABILIZERCONTROL UNIT

2324

AH-800 AHRU NO.1

ARINC 429OUT #1

ARINC 429OUT #2

(H)(L)

(H)(L)

G7G8

E5E6

83

84

2930

C1J1B-E5C1J1B-E5FIG

2-11-6

9J1

1J1B

PITCH WHEEL (+)PITCH WHEEL (-)

TURN KNOB+15V INPUT

H

UV

A

B

C


+15 V MODE POWER

AP ENG PUSHBUTTON

YD ENG PUSHBUTTON

CPL PUSHBUTTONAP ENG LAMP

YD ENG LAMP

1023

24

25CPL L LAMP

21


60

73

59

75

6869

129J1

3334

C190J2A-83C190J2A-84

85

8355

8684

99

C190J2A-43C190J2A-44

FIG2-11-6

AP ENGAGE SELECT PB IN

YD ENGAGE SELECT PB IN

COUPLE PB INPUTAP ENGAGE ANNUNC OUTYD ENGAGE ANNUNC OUT

COUPLE ANNUNC OUT

(H)(L)

ARINC 429CH.3 IN

(H)

(L)

ARINC 429CH.11 IN

(H)(L)

ARINC 429CH.4 IN

(H)(L)

ARINC 429CH.0 IN

IC-600 IAC NO.1

101102

88

SM-200 AILERONSERVO

JF

B

A

12J1

99

100

LINKAGEAILERON

35

3643

44

190J2A

SM-200 ELEVATORSERVO

JF

B

A

12J1

LINKAGEELEVATOR

105

10689103

SM-200 RUDDERSERVO

JFB

A

14J1

LINKAGERUDDER

104

NL

989787

29

90

C9J1-60

C9J1-61

190J2B

FIG 2-11-6

(+)(-)

TACHOMETEROUTPUT

(+)(-)

CLUTCHEXCITATION

(+)(-)

MOTOR INPUT(CCW DRUMROTATION)

N

LTACHOMETEROUTPUT

(+)(-)

MOTORINPUT

(+)(-)

CLUTCHEXCITATION

DC-550 DISPLAYCONTROLLER NO.1

32

33

27

115J1

28V AP CLUTCH

28V AUTOPILOTDISCONNECT

TCS INPUT

COPILOTS GA SWITCH

PLT SERVOPOWER

PILOTS WHEELMASTER SWITCH

COPILOTS WHEELMASTER SWITCH

PILOTS TCS COPILOTS TCS

AUTOPILOTSERVO POWER

AD-51314-R1@

190J2B

190J2A

76

70

AP DISCONNECT28 V IN

TCS 28 V INPUT

GA INPUT

(H)(L)

ARINC 429CH.1 IN

(H)(L)

ARINC 429CH.10 IN

(H)(L)ELEVATOR SERVO OUT

(H)(L)

ELEVATOR TACHFEEDBACK IN

28 V AP CLUTCH SERVO

(H)(L)

RUDDER SERVO OUT

28 V YD CLUTCH

(+)(-)

MOTOR INPUT(CCW DRUMROTATION)

(+)(-)

CLUTCHEXCITATION

N

L (+)(-)

TACHOMETEROUTPUT

(H)(L)

AILERONSERVO OUT

28 V APCLUTCH

(H)(L)

AILERON SERVOTACH FEEDBACK IN

96

95

RUDDER SERVOTACH FB IN -

RUDDER SERVOTACH FB IN +-15 V EXTERNAL OUT

TURN KNOB INPUT

+15 V EXTERNAL OUT

TURN KNOB OUT OF DETENT

PITCH WHEEL IN +

PITCH WHEEL IN -

190J2B

190J2A

190J2B

PILOTS GA SWITCH

TOC190J2B-70

C115J1-32

TURN KNOB-15 INPUT

TURN KNOB SIGNAL

TURN KNOB NOTIN DETENT

6061

11J1

11J1

FIG2-11-6

190J2B

C1J1B-E6C1J1B-E6

190J1B

TOHSCU

Figure 2-11-5. Pilot’s Autopilot/Yaw Damper Interface



10

36

GC-550 GUIDANCECONTROL UNIT

CPL PUSHBUTTON

CPL R LAMP

+15V MODE POWER21

11J1

61

64

AZ-850 MADC NO. 2

ARINC 429BUS NO. 2

C9J1

63

ARINC 429BUS NO. 2

60

TO HORIZONTALSTABILIZERCONTROL UNIT

FIG 2-11-59J1-60

9J1-61 4443

G8

E6

AH-800 AHRU NO. 2

ARINC 429BUS NO. 2

C9J1B

E5

ARINC 429BUS NO. 2

G7

190J2A-83190J2A-84 FIG 2-11-5

8483

LH

LH

LH

LH

C190J2A B

IC-600 IAC NO.2

55

99

23

24

3029

ARINC 429CH 3 INL

H

ARINC 429CH 10 INL

H

ARINC 429CH 4 IN(PRI AHRS)L

H

ARINC 429CH 11 IN(SEC AHRS)L

H

75TURN KNOBOUT OF DETENT

H


TURN KNOB NOTIN DETENT

CTURN KNOB + 15V INPUT

COUPLE PB INPUT

COUPLE ANNUN OUT

129J1

129J1

C129J1B DC-550-DISPLAYCONTROLLER NO. 2

TCS

C115J1

3270TCS 28 V INPUT

28V/OPENFROM PILOTS/COPILOTSTCS SWITCHES

FIG 2-11-5

AD-51315-R1@

190J2A-35190J2A-43

190J2A-44190J2A-36

FIG-2-11-5

FIG 2-11-51J1B-E51J1B-E6

Figure 2-11-6. Copilot’s Autopilot/Yaw Damper Interface



B. Autopilot/Yaw Damper Engage Logic

The PRIMUS® 1000 autopilot/yaw damper is a fail-passive system that requires a singlevalid AHRS heading gyro and two valid AHRS attitude sources for autopilot and yawdamper engagement (see Figure 2-11-7). Dual AH-800 AHRS provide the required verticaland directional gyro requirements in this aircraft. The pilot’s AHRS is the primary sourcefor the autopilot/yaw damper and monitor. The copilot’s AHRS is used as the reference forcomparison monitoring functions.

Autopilot and yaw damper engagement is controlled from the GC-550 Guidance ControlUnit. The AP quick disconnect switch, electric trim switch and touch control steering (TCS)switch both located on the control wheels will also affect autopilot engagement, as will thesecondary pitch trim switch located on the pedestal.

Pressing the AP ENGAGE select switch engages the autopilot and yaw damper if allengage logic is valid. Pressing the YD ENGAGE select switch engages only the yawdamper. Pressing the AP ENGAGE select switch when the autopilot is engaged,disengages the autopilot only. Pressing the YD ENGAGE select switch with the autopilotand/or the yaw damper engaged will disengage both.

Autopilot pitch trim is engaged and disengaged as a function of autopilotengage/disengage.

(1) Yaw Damper Engagement

To engage the rudder servo clutch, AND gate 1 must have a high output. To satisfythis requirement, latch 2 must have its Q output high and AND gate 3 must have ahigh output. See Figure 2-11-7.

Latch 2 will go high as a function of pressing the YD ENGAGE button, or the APENGAGE button on the GC-550 Guidance Control Unit.

AND gate 3 will go high as a function of:

• Yaw rate valid from both pilot and copilot’s AHRS

• Stick shaker is not active

• Internal monitor discretes are valid

• AHRS valid from both pilot and copilot’s AHRS

• AP DISC switch is not active

• Servo power is NOT less than 10 V for more than 0.25 sec

• IC Bus is valid and the cross-side SG is in reversion

• Aircraft ID pins are correctly configured.



(2) Autopilot Engagement

To engage the aileron and elevator servo clutches, AND gate 2 must have a highoutput. To accomplish this, latch 1 must have its Q output high, TCS must not beactive and AND gate 4 must have a high output. See Figure 2-11-7.

Latch 1 will go high as a function of pressing the AP ENGAGE button on the GC-550Guidance Control Unit. AND gate 4 will go high as a function of:

• Yaw rate valid from both pilot and copilot’s AHRS

• Stick shaker is not active

• Internal monitor discretes are valid

• AHRS valid from both pilot and copilot’s AHRS

• AP DISC switch is not active

• Servo power is NOT less than 10 V for more than 0.25 sec

• IC Bus is valid and the cross-side SG is in reversion

• Aircraft ID pins are correctly configured

• Pilot or copilot electric trim or secondary trim is NOT active

• TURN knob is in detent position

• Windshear detect discrete is not enabled.



INTERNALMONITORDISCRETES(NOTE 1)

A

A/C ID VALID

AP ENG

WINDSHEARDETECT

7372

74

75

190J2A

190J2B

1324

AP ENG ANN

YD ENG ANN84

86

83

85

H

129J1TKOD

DETENT

IN

OUT

B

J1190J2B-73

+28V


GPWS/WINDSHEARCOMPUTER

LTGCOMMON

24

YD ANN

AP ANN

23

YD ENG SEL

9

AP ENG SEL

8

21

GC-550 GUIDANCECONTROL UNIT

49

45

DN

UP28V

3334

AH-800 NO. 1E5E6

9J1B

8384

AH-800 NO. 2E5E6

C9J1B

TOC190J2A83, 84

190J2A

YAW RATE NO. 1 VALID

AHRS NO. 1 VALID

YAW RATE NO. 2 VALID

AHRS NO. 2 VALID

AHRSVALID

YAW RATEVALID

70

190J2B

SECONDARYTRIM SWITCH

67

66

DN

UP

COPILOT'S ELECTRIM SWITCH

190J1A

76

77

DN

UP

PILOT'S ELECTRIM SWITCH

59

190J2B

+15V129J1-C

PILOT'SA/P DISC

COPILOT'SA/P DISC

C190J2B-76C190J2B-90C115J1-33

190J2B

3,4

90

1,2

76

190J2B

X4 X4

115J1-33

SERVO POWERMONITOR

AP DISC

<10V 0.25 SEC

SECONDARY TRIM

ACTIVE

ACTIVE

PILOTMANUAL TRIM

ACTIVE

COPILOTMANUAL TRIM

SECONDARY TRIM

ACTIVE

MANUALTRIM

ACTIVE

TRIM

ACTIVE

YD ENG

IC-600 IAC NO. 25758

C190J2A

5758

190J2A

COPILOT'STCS SWITCH

PILOT'STCS SWITCH

ICB VALID

IC-600 IAC NO. 1

A

28VDC

A/P SERVO CB

34

T

R

QAP ENGAGE

1

AP ENABLE

TCS

TO PFD FORENGAGE/DISENGAGE

ANNUNCIATION

T

R

QYD ENGAGE

2

YD ENABLE

TO PFD

87

88

89

SERVOCLUTCH

F J

SERVOCLUTCH

F J

HSCU

115J1-27FOR AURAL ANNUNCIATION

SERVOCLUTCH

F J

ELEVATOR

AILERON

RUDDER

AD-51316-R2@

NOTES: 1.

2.

THE FOLLOWING INTERNALGENERATED LOGIC WILLDISENGAGE THE AP/YD: AP DISCONNECT MONITOR FAIL YD DISCONNECT MONITOR FAIL AP PREFLIGHT TEST FAIL PITCH ROLL OR YAW SERVO MODEL FAIL EXCESSIVE ROLL RATE EXCESSIVE VERTICAL ACCEL MEMORY TEST FAIL TRIM RUN AWAY FAILURE AHRS MISCOMPARE

ALL LOGIC GATES SHOWNREPRESENT SIMPLIFIED INTERNALSOFTWARE IMPLEMENTATION, HOWEVER ALL SOFTWARE LOGICIS NOT SHOWN.

11J1

78STALL WARNING 28VDC DISCRETE

190J2B

190J1B

190J2B12J1 12J1

13J113J1

14J1 14J1

FOR ENGAGE/DISENGAGEANNUNCIATION

63

115J1-33

X-SIDESG REV

4

2

1

Figure 2-11-7. Autopilot/Yaw Damper Engage Logic



C. Roll Axis Autopilot Servo Loop

The autopilot roll axis servo loop (See Figure 2-11-8 and Table 2-11-5) is the same for alllateral steering modes. Since there is only one autopilot system and one servo in the rollaxis, it does not matter if the steering command is heading select, or localizer, the path tothe servo is the same. The autopilot control loop is connected in parallel to the aircraft’sprimary control rigging through cables.

The autopilot roll axis algorithim includes some compensation for possible deadzone in theaileron control system. Slight control column activity may be observed in some flightconditions, since the control law is designed to minimize the autopilot reaction to controlsystem deadzone.

(1) SM-200 Roll Servo Drive and Bracket

The SM-200 translates electrical input signals into a clutched mechanical output.This output is used to drive the elevator in response to pitch axis autopilotcommands. A description of servo functions follows:

(a) Servo Clutch

The servo clutch is engaged as a function of autopilot engagement, or releaseof touch control steering (TCS) if the autopilot was previously engaged. Theclutch will disengage anytime the autopilot is disconnected, or the TCS switchis pushed and held.

(b) Servo Torque Motor

The servo torque motor receives dc current from the IC-600 IAC pitch axis.With the servo clutch engaged, the torque motor output drives a power geartrain through mechanical coupling. The gear train output in turn provides thedrive to move the elevator to the desired position. With the autopilot notengaged, any input drive to the servo motor will not be coupled to the elevator.

(c) DC Tach Generator

The DC tach generator is mechanically connected to the servo torque motorand will provide an output back to the IC-600 IAC anytime the servo torquemotor drives. The DC tach generator provides two functions:

Provides an elevator rate of travel signal to the IC-600 IAC. This signal isused as a damping term. When the elevator is commanded to a position, itshould move to that position smoothly and stop still, not move or hunt aboutthat position.

In the IC-600 IAC the DC tach generator signal is also integrated to deriveelevator position feedback. This signal is used to ensure that the elevatortorque motor has driven the elevator as properly commanded.



(2) IC-600 Integrated Avionics Computer (IAC)

The IC-600 IAC receives sensor data and command inputs and processes this datain accordance with any vertical steering mode that is active. Since this is a digitalcomputer, this processing is accomplished through software. In order to provide acurrent to drive the servo torque motor, this digitally processed signal must bechanged into analog form.

Additionally, to insure safe operation, certain functions and values of certainparameters are monitored in the IC-600 IAC to ensure that the autopilot will beautomatically disconnected should a safety critical malfunction occur. A descriptionof the IC-600 pitch axis autopilot servo drive follows:

(a) Roll Axis Attitude Loop

The roll axis attitude loop processes roll attitude and roll rate of change fromthe pilot’s AHRS to derive a corresponding roll rate term. Both roll attitudeand roll rate terms are gain programmed as a function of IAS.

The flight director roll command is limited to ±35° and rate limited to 7° persecond by the autopilot. The rate limited roll attitude reference command isused as an input to the aileron servo loop.

(b) Aileron Servo Loop

The aileron servo loop uses the roll attitude loop command to compute anaileron servo pulse width command with the autopilot engaged. If the autopilotis not engaged, the aileron servo pulse width is zero.

Aileron servo position is derived by integrating the aileron servo DC tachgenerator feedback signal. The DC tach generator signal is also used as adamping term in positioning the ailerons. Aileron servo current is passedthrough a current limiter and servo driver before being sent to the servoamplifier.

(c) Current Limiter

Current limiting is performed on the servo command signal to insure that theproper servo drive values are established.

(d) Servo Amplifier

The servo amplifier acts as a switch to provide drive current to the aileronservo torque motor. A servo enable discrete is applied as a function ofautopilot engagement. The servo requires 1-ampere current drive capability.The servo amplifier supplies a 480 Hz pulse-width-modulated 28-volt bipolaroutput. The pulse width command output is compared with a 480 Hz sawtooth signal to generate the pulse width control for the servo driver. The servoloop software executes at 240 Hz so that the servo amplifier output is thesame for two complete duty cycles.



The primary processor provides a discrete output which enables the aileronservo amplifier. If this discrete is not available, the servo amplifier is forced toa zero duty cycle. Also a latched heartbeat monitor and a latched powersupply monitor (both not shown), are required to enable the servo amplifierdriver.

For the primary processor servo amplifier enable to be active:

• All monitors must be valid

• Both attitude sources (AHRS) must be valid

• Communications with the secondary processor must be valid

• All internal processor valids must be valid.

Table 2-11-5. Autopilot Roll Axis Operating Limits


Autopilot MAX Bank Limit ± 35°

MAX Rate Limit 7.0°/sec

Limit AfterEngagement

± 35° If the aircraft is rolled to anangle > 35° using TCS, theautopilot will roll the aircraft to35° and maintain.

Heading Hold Bank Angle less than 3° plus 3 seconds

Roll Hold Bank Angle The bank angle is held if thebank is greater than 6° but lessthan 35° and the bank wasinitiated using TCS.

Turn Knob Bank angle Max bank angle is 35°Max roll rate is 3°/sec.

Pitch Wheel Pitch Angle ±20°



(Blank Page)



SERVOLOOPGAIN

PULSEWIDTH

COMMANDLIMIT AND

D/A

PROCESSORCONTROL

101

102

99

100

8828V DC

A/D

TACHGAIN

POSITIONGAIN

AP

190J2B 12J1

C

C

F

J

SM-200 AILERON SERVO DRIVE AND BRACKET

M

T

TORQUEMOTOR

DC TACHGEN

POWERGEARTRAIN

CLUTCH

DRUM

LINKAGE AILERON

IC-600 IAC (NO. 1)

AD-51317-R1@

+

-

SERVO ENABLEDISCRETE

TACH

AIRCRAFTRESPONSE

+

-

POSITIONINTEGRATOR

AUTOPILOTCMD

CURRENTLIMITER

SERVODRIVER/

AMPLB

A

N

L

Figure 2-11-8. Autopilot Roll Axis Servo Loop



D. Pitch Axis Autopilot Servo Loop

The pitch autopilot servo loop (see Figure 2-11-9 and Table 2-11-6) is the same for allvertical steering modes. Since there is only one autopilot system and one servo in thepitch axis, it does not matter if the steering command is altitude hold or vertical speed hold,the path to the servo is the same. The autopilot servo loop is connected in parallel to theaircraft’s primary control rigging through cables.

For the servo to perform its job it receives a drive signal from the IC-600 IAC. This can bethought of as the autopilot command to drive the elevator. As the servo drives, it movesthe control rigging which in turn moves the elevator. As the elevator moves, a feedbacksignal from the servo is sent to the IC-600 IAC. This signal can be thought of as autopilotresponse. When the autopilot response equals the autopilot command the elevator stopsmoving. As the autopilot command is satisfied, the command gets smaller in size and thefeedback signal will now drive the elevator back to its starting position.

(1) SM-200 Elevator Servo Drive and Bracket

The SM-200 translates electrical input signals into a clutched mechanical output.This output is used to drive the elevator in response to pitch axis autopilotcommmands. A description of servo functions follows.

(a) Servo Clutch

The servo clutch is engaged as a function of autopilot engagement, or releaseof touch control steering (TCS) if the autopilot was previously engaged. Theclutch will disengage anytime the autopilot is disconnected, or the TCS switchis pushed and held.


The servo torque motor receives dc current from the IC-600 IAC pitch axis.With the servo clutch engaged, the torque motor output drives a power geartrain through mechanical coupling. The gear train output in turn provides thedrive to move the elevator to the desired position. With the autopilot notengaged, any input drive to the servo motor will not be coupled to the elevator.



• Provides an elevator rate of travel signal to the IC-600 IAC. This signal isused as a damping term. When the elevator is commanded to a position,it should move to that position smoothly and stop still, not move or huntabout that position.

• In the IC-600 IAC the DC tach generator signal is also integrated to deriveelevator position feedback. This signal is used to ensure that the elevatortorque motor has driven the elevator as properly commanded.




The IC-600 IAC receives sensor data and command inputs and processes this datain accordance with any vertical steering mode that is active. Since this is a digitalcomputer, this processing is accomplished through software. In order to provide acurrent to drive the servo torque motor, this digitally processed signal must bechanged into analog form.

Additionally, to insure safe operation, certain functions and values of certainparameters are monitored in the IC-600 IAC to ensure that the autopilot will beautomatically disconnected should a safety critical malfunction occur. A descriptionof the IC-600 pitch axis autopilot servo drive follows:

(a) Servo Amplifier

The servo amplifier acts as a switch to provide drive current to the servotorque motor. A servo enable discrete is applied as a function of autopilotengagement. The servo requires 1-ampere current drive capability. The servoamplifier supplies a 480 Hz pulse-width-modulated 28-volt bipolar output. Thepulse width command output is compared with a 480 Hz saw tooth signal togenerate the pulse width control for the servo driver. The servo loop softwareexecutes at 240 Hz so that the servo amplifier output is the same for twocomplete duty cycles.

The primary processor provides a discrete output which enables the elevatorservo amplifier. If this discrete is not available, the servo amplifier is forced toa zero duty cycle. Also a latched heartbeat monitor and a latched powersupply monitor (both not shown), are required to enable the servo amplifierdriver.

For the primary processor servo amplifier enable to be active:

• All monitors must be valid

• Both attitude sources must be valid

• Communications with the secondary processor must be valid

• All internal processor valids must be valid.

(b) Tach Integrator

The servo tach generator provides an elevator rate of travel signal as adamping term. This same rate signal is integrated to derive elevator servoposition feedback. If rate is integrated over time, distance travelled, orposition is derived.

(c) Current Limiter




(d) Servo Loop Processing

The servo position command is summed with the servo position feedback andrate signals. Servo loop gain takes into account the mechanical advantage ofthe cable rigging in the pitch axis.

(e) Pitch Trim Threshold Sensor

The pitch axis of the autopilot incorporates a trim function to compensate theaxis for long term, steady state loads felt on the elevator. These loads can besuch things as fuel burn and passenger movement. Rather than use elevatorservo current to hold the elevator in position for these long term states, it isdesirable to electrically unload the servo and mechanical drive to repositionthe elevator aerodynamically and hold the elevator in this new position. Toaccomplish this, the autopilot monitors the amount of current going to theelevator servo.

(f) Pitch Wheel

Moving the pitch wheel will cause a rate generator output. The direction of thepitch wheel motion will determine the polarity of the output, while the speed ofpitch wheel motion will determine the amplitude of the signal. The pitch wheelsignal is then TAS gain programmed then summed with the autopilotcommand and aircraft response.

Table 2-11-6. Pitch Channel Axis Operating Limits


AP Pitch Limit ± 20°

TCS Pitch Limit ± 20°

Pitch Wheel ± 20°



E. Pitch Axis Autopilot Trim

The autopilot processor performs elevator trim control based on elevator servo currentdemand. Elevator trim engagement is controlled by the autopilot engage logic and the trimclutch is wired to the same clutch output as the aileron and elevator servo clutches.

The trim actuator is driven in a manner to move the horizontal stabilizer to reduce the airload on the primary elevator servo. The trim drive is a 0.625 Hz pulse width modulated 28V dc output. The pulse width is controlled as a function of airspeed.

Elevator trim rate is programmed with TAS to provide variable trim rate based on flightconditions. The use of TAS for gain programming allows the long term elevator response(trim rate) to decrease with increasing altitude.

When elevator servo current exceeds a predetermined threshold for a given period of time,this is considered to be a steady state error and trim will run. Out of the threshold detectorand delay will be a positive or negative current. This will turn on the up or down sensoraccordingly and allow the trim motor to run. As the trim runs, the horizontal stabilizer isre-positioned and the air load on the elevator primary servo is reduced. When this loadfalls below the threshold level, trim stops running.

When the up/down sensor allows an output, an internal clock runs for 10 seconds. If a trimmalfunction has occurred and the trim system has not sufficiently reduced elevator servocurrent, the other trim threshold sensor will allow an output to annunciate an out of trimcondition CAS message. This annunciation informs the pilot that the aircraft is out of trimand that a disconnect while out of trim may cause a slight maneuver. The mistrimannunciation will disappear when the aircraft finally reaches a trim state or when theautopilot is disconnected.

To manually trim the aircraft the pilot would take control of the aircraft, disengage theautopilot and retrim the aircraft.



TRIMTHRESHOLDSENSOR AND

10 SEC TIME DELAY

SERVOLOOPGAIN

AUTOPILOTCMD

POSITIONGAIN

AIRCRAFTRESPONSE

TACH

PULSEWIDTH

COMMANDLIMIT AND

D/A

PROCESSORCONTROL

TACHGAIN

AD

CURRENTLIMITER

TACH FEEDBACK

SERVODRIVER/

AMPL

SERVO ENABLEDISCRETE

28V AP CLUTCH

SERVO

UP/DOWNSENSOR

TASPROGRAMMING

AP ENG

TIME DELAYTRIMTHRESHOLD

AP PITCHTRIM

MOTORDRIVER

PULSE WIDTH MODULATORBASED ON SPEED

AND SERVO CURRENT

IC-600 IAC NO. 1

92

93

TO FWCMISTRIM MSG ON EICAS

UPTRIM

DOWNTRIM

(+) RETRACT

(-) EXTEND

190J2B

6869

PC-400 TURN PITCHCONTROLLER

U

V

(+) PITCH WHEEL

(-) PITCH WHEEL

A

B

C

L

N

C

F

SM-200 ELEVATOR SERVO DRIVE AND BRACKET

M

TORQUEMOTOR

CLUTCH

POWERGEARTRAIN

T

LINKAGEELEVATOR

DRUM

J

AD-51318-R1@

105106

103

104

89

+

-

+

+

-

(NON HONEYWELL)HORIZONTAL STABILIZER

CONTROL UNIT

13J1

Figure 2-11-9. Pitch Autopilot Servo Loop



F. Yaw Damper Rudder Axis Servo Loop

The yaw damper servo loop (see Figure 2-11-10) is independent from the autopilot in that itcan be engaged, without the autopilot. The yaw damper servo is connected in parallel withthe aircraft’s primary rigging through cables. The yaw damper provides yaw damping fortransient wind gust conditions.


The IC-600 IAC receives sensor data and command inputs and processes this data.Since this is a digital computer, this processing is accomplished through software. Inorder to provide a current to drive the servo torque motor, this digitally processedsignal must be changed into analog form.

Additionally, to insure safe operation, certain functions and values of certainparameters are monitored in the IC-600 IAC to ensure that the yaw damper will beautomatically disconnected should a safety critical malfunction occur. A descriptionof the IC-600 IAC rudder axis servo drive follows:

(a) Servo Amplifier

The servo amplifier acts as a switch to provide drive current to the servotorque motor. Servo enable is applied as a function of yaw damperengagement. The servo requires 1-ampere current drive capability. The servoamplifier supplies a 480 Hz pulse width modulated 28V bipolar output. Thepulse width command output is compared with a 480 Hz saw tooth signal togenerate the pulse width control for the servo driver. The servo loop softwareexecutes at 240 Hz so that the servo amplifier output is the same for twocomplete duty cycles.

(b) Tach Integrator

Position feedback is achieved taking the rudder servo tachometer signal whichis a rate of travel term and integrating it. Integrating rate of travel over time,derives distance travelled, or position.

(c) Current Limiter


(d) Rudder Washout

The yaw damper is primarily responsive to short term rate signals. Since theservo loop is an electrical-mechanical operating system, it is possible that theyaw rate attitude loop command and servo feedback signals do not cancel. Toeliminate this stand-off condition, any long term output of the servo amplifier isfed to the rudder washout integrator. The output of the integrator is invertedand sent back into the servo loop to eliminate the stand-off condition.



Rudder washout integrates any steady state force and subtracts this from theposition reference. This washout allows the pilot to manually re-trim therudder with the yaw damper engaged. The washout also eliminates anysteady state forces on the rudder.

(2) SM-200 Servo Drive and Bracket

The SM-200 translates electrical input signals into a clutched mechanical output.This output is used to drive the rudder in response to yaw axis commands. Adescription of servo functions follows:

(a) Servo Clutch

The servo clutch is engaged as a function of autopilot or yaw damperengagement. The clutch will disengage anytime the yaw damper isdisconnected.


The servo torque motor receives dc current from the IC-600 IAC yaw axis.With the servo clutch engaged, the torque motor output drives a power geartrain through mechanical coupling. The gear train output in turn provides thedrive to move the rudder to the desired position. With the yaw damper notengaged, any input drive to the servo motor will not be coupled to the rudder.



• Provides a rate of travel feedback signal to the IC-600 IAC. This signal isused as a damping term. When the rudder is commanded to a position, itshould move to that position smoothly and stop still, not move or huntabout that position.

• In the IC-600 IAC the DC tach generator signal is also integrated to deriverudder position feedback. This signal is used to ensure that the ruddertorque motor has driven the rudder as properly commanded.



YAW RATE D/ASERVOLOOPGAIN

PULSEWIDTH

COMMANDLIMITER

PROCESSORCONTROL

A/D

97

98

8728V DCYD CLUTCH

A/D

TACHGAIN

POSITIONFEEDBACK

LIMITER

190J2B 14J1

A

B

L

N

K

F

J

SM-200 RUDDER SERVO DRIVE AND BRACKET

M

T

TORQUEMOTOR

DC TACHGEN

POWERGEARTRAIN

CLUTCH

DRUM

LINKAGE RUDDER

IC-600 IAC (NO. 1)

AD-51319-R1@

+

-

+ +

-

SERVOENABLE

95

96

+

ROLL ATTITUDEPROGRAMMERROLL ATTITUDE

TAS

RUDDERWASHOUT

-SERVO

DRIVER/AMPL

CURRENTLIMITER

Figure 2-11-10. Yaw Damper Servo Loop



4. Fault Monitoring

A. Autopilot/Yaw Damper Monitoring Overview

The autopilot/yaw damper monitoring function is provided by the IC-600 IAC primaryprocessor. The primary processor utilizes independent hardware for sensor input andservo feedback data used by the monitoring function. This separation ensures thathardware failures will not affect the autopilot control function.

The primary processor provides dedicated disconnect hardware for the monitoring function.This provides the monitors with the capability of disengaging the autopilot and yaw damperindependent of the autopilot processor hardware. All automatic disconnects which resultfrom monitor trips store an event code in non-volatile memory for subsequent recall duringground maintenance testing.

The disengage path is tested during autopilot power-up to ensure that latent failures do notinhibit monitor operation. The servo amplifier disable path (heartbeat monitor, powersupply monitor, monitor processor valid) are all individually tested at power-up. Thesetests consist of driving the pitch, roll and yaw servos; and validating proper tachometerfeedback and current sensing.

The monitors use computed TAS and IAS to model gain programmers and the pitch g limit.If the air data is not valid, the following default values are used:

• TAS = 300 kts

• IAS = 220 kts.

B. Hardover Malfunction Protection

The distributed processor architecture of the PRIMUS® 1000 system is designed to preventhazards such as autopilot hardovers. This protection is provided by ensuring that failuresin either processor will be defeated or minimized by the monitors and/or limiters in theother processor. In general, failures of the primary processor (flight director), will not resultin an abrupt aircraft response because the attitude command path is limited to a normalcontrol response envelope in the secondary processor (autopilot). This is insured by a±25° magnitude limit and a ±0.3g rate limit in the pitch axis and a ±35° magnitude limit anda ±7°/sec rate limit in the roll axis in the autopilot processor. Failures of the autopilotprocessor are detected by the monitors in the primary processor and result in an autopi-lot/yaw damper disconnect, well in advance of exceeding autopilot hazard criteria.

C. System Response to Failures

In the event of an autopilot processor fault, the monitor will disconnect the autopilot prior toany significant aircraft response. Upon monitor disconnect, a red AP will be displayed onthe PFD.



D. Monitor Description

(1) Pitch Servo Position Monitor

The pitch servo position monitor models the pitch attitude loop and servo loop andgenerates a predicted elevator servo position. The predicted servo position iscompared against actual position feedback. The actual servo feedback is generatedby lagging the servo tachometer feedback in the monitor process. The monitor triplevel is programmed with IAS to allow a somewhat uniform aircraft responseexposure. Any failure which results in exceeding the monitor trip level for a period of0.3 seconds will result in an automatic disconnect of the autopilot and yaw damper.

(2) Primary Pitch Attitude Comparison

The primary pitch attitude comparison monitor compares the on-side primary pitchattitude used in the primary processor with the pitch attitude used in the secondaryprocessor. This monitor validates the hardware path between processors andassures that the pitch attitude computations of both processors agree.

A trip of the pitch comparator monitor will prevent autopilot and yaw damperengagement and will cause a latched disengage if the autopilot/yaw damper wereengaged at the time of the monitor trip.

If the pitch comparator exceeds a 2° threshold, the monitor will trip.

(3) Secondary Pitch Attitude Comparision Monitor

The secondary pitch attitude comparison monitor compares the on-side primary pitchattitude with the on-side secondary pitch attitude. This monitor validates the attitudesource used by the pitch servo position monitor.

(4) Normal Accelertion Monitor

The normal acceleration monitor provides an additional means of detecting autopilotmalfunctions. The sensor input for this monitor is provided by the AHRS. Themonitor disconnects the autopilot and yaw damper if normal acceleration changes bymore than ±0.6 g’s for more than 0.4 seconds.

(5) Roll Servo Position Monitor

The roll servo position monitor models the roll attitude loop and servo loop andgenerates a predicted aileron servo position. The predicted servo position iscompared against actual position feedback. The actual servo feedback is generatedby lagging the servo tachometer feedback in the monitor process. The monitor triplevel is programmed with IAS to allow a somewhat uniform aircraft responseexposure. Any failure which results in exceeding the monitor trip level for a period of0.3 seconds will result in an automatic disconnect of the autopilot and yaw damper.



(6) Primary Roll Attitude Comparison Monitor

The primary roll attitude comparison monitor compares the on-side primary rollattitude used in the primary processor with the roll attitude used in the secondaryprocessor. This monitor validates the hardware path between processors andassures that the roll attitude computations of both processors agree.

A trip of the roll comparator monitor will prevent autopilot and yaw damper engage-ment and will cause a latched disengage if the autopilot/yaw damper were engagedat the time of the monitor trip.

If the roll comparator exceeds a 2° threshold, the monitor will trip.

(7) Secondary Roll Attitude Comparison Monitor

The secondary roll attitude comparison monitor compares the on-side primary rollattitude with the on-side secondary roll attitude. This monitor validates the attitudesource used by the roll servo position monitor.

(8) Roll Rate Monitor

The roll rate monitor provides an additional means of detecting autopilot malfunc-tions. The monitor disconnects the autopilot and yaw damper if actual roll rateexceeds 12°/sec for more than 0.5 seconds.

(9) Yaw Servo Position Monitor

The yaw servo position monitor models the yaw rate loop and servo loop andgenerates a predicted elevator servo position. The predicted servo position iscompared against actual position feedback. The actual servo feedback is generatedby lagging the servo tachometer feedback in the monitor process. The monitor triplevel is programmed with IAS to allow a somewhat uniform aircraft responseexposure. Any failure which results in exceeding the monitor trip level for a period of0.6 seconds will result in an automatic disconnect of the autopilot and yaw damper.

(10) Auto Trim Runaway Monitor

The auto trim runaway monitor detects any condition which results in the autopilotprocessor commanding trim while the elevator servo current does not indicate a needfor trim. The auto trim runaway monitor will disconnect the autopilot and yawdamper immediately upon detecting a trim runaway condition.

(11) Auto Trim Inoperative Monitor

The trim inoperative monitor provides a warn annunciator in view of the pilot (CASmessage) to indicate that the elevator is not properly trimmed. This monitor doesnot disconnect the autopilot.



(12) Autopilot/Yaw Damper Disconnect Monitor

The autopilot and yaw damper disconnect monitor detect a failure of the system todisengage the autopilot and yaw damper in response to the autopilot disconnectswitch being pressed. This monitor ensures the disconnect discrete and AP and YDengage status are valid. If self engagement within 0.6 seconds of a disconnect isdetected, the processor outputs an invalid state on the servo amplifier drive enable.This action prevents the IAC from applying any torque to the autopilot and yawdamper servos.
