bell 407 - flight manual

•

407 - Flight Manual 407 Flight Manual 1 - Temporary Revisions

Log of Temporary Revisions (16 Nov 2006)

TR-9 - Sustained Hover and Vertical Takeoff/Landing Operations With Tailwind (15 Jan 2002)

TR-10 - Incorporation of Oil Cooler Blower Inlet Ducts and Bearing Airflow Shields (Rev 1 - 25 Jul 2002)

TR-11- FADEC Software Version 5.356 (16 Nov 2006)

407-FM-1 - Flight Manual (Revision 7 - 30 Jul 2008)

Front Matter

Section 1 - Limitations

Section 2 - Normal Procedures

Section 3 - Emergency/Malfunction Procedures

Section 4 - Performance

Section 5 - Weight and Balance

Appendix A - Optional Equipment Supplements

407-FM-1-ATO - Title Page for Republic of Philippines Registered Helicopters (Basic Issue: 12 Nov 1996)

407-FM-CTA - Title Page for Brazil Registered Helicopters (Basic Issue: 03 Apr 1998)

407-FM-IAC AR - Title Page for Interstate Aviation Committee - Aviation Register Commonwealth of Independent States (Basic Issue: 20 May 1999)

407 - Flight Manual Supplement

407-FMS-1 - Lightweight Emergency Flotation Landing Gear (Basic Issue: 11 Apr 1996)

407-FMS-2 - High Skid Gear (Basic Issue: 14 Feb 1996)

407-FMS-3 - Particle Separator (Reissue - 16 Dec 2002)

407-FMS-4 - Snow Deflector (Reissue - 16 Dec 2002)

407-FMS-5 - Cargo Hook (Revision 1 - 25 Mar 2008)

407-FMS-6 - Auxiliary Fuel Kit (Basic Issue: 20 Mar 1996)

407-FMS-7 - Litter Kit (Revision 1 - 16 Sep 1999)

407-FMS-17 - Cargo Tiedown Provisions Kit (Basic Issue: 01 Apr 1996)

407-FMS-20 - KLN 89B GPS Navigator (Revision 1 - 26 Nov 1996)

407-FMS-21 - Fire Detection System (Reissue - 07 Jul 2004)

http://127.0.0.1:1584/index.php?action=getpdf&idfile=2118�



























407-FMS-22 - Auxiliary Vertical Fin Strobe Lights (Basic Issue: 10 May 1996)

407-FMS-23 - RYAN Traffic Collision Avoidance Device (Basic Issue: 15 May 1996)

407-FMS-25 - Quiet Cruise Mode (Revision 1 - 30 Jul 2008)

407-FMS-28 - Increased Internal Gross Weight (Reissue - 16 Dec 2002)

407-FMS-31 - Increased APC Starter Generator Load Kit (Basic Issue: 15 Jun 2005)

407-FMS-CAA - United Kingdom Registered Helicopters (Basic Issue: 08 Jan 2002)

407-FMS-IAC AR - Interstate Aviation Committee - Aviation Register Commonwealth of Independent States (Basic Issue: 20 May 1999)

407 - Manufacturer's Data

407-MD-1 - Manufacturer's Data (Revision 4 - 30 Apr 2008)

Front Matter

Section 1 - Systems Description

Section 2 - Handling and Servicing

Section 3 - Conversion Charts and Tables

Section 4 - Expanded Performance














LOG OF TEMPORARY REVISIONS

This Log of Temporary Revisions provides the current status of each Temporary Revision issued againstthe basic Flight Manual. It should be inserted at the back of the Flight Manual binder for quick and easyreference.

TEMP. REV. NO. TITLE DATE ISSUED DATE CANCELED

BHT-407-FM-1 FADEC Fault Annunciation Interpretation

Revision 103 December 1996

15 June 2000

BHT-407-FM-1 NP Overspeed Trip Increase 03 December 1996 15 June 2000

BHT-407-FM-1 Airspeed Change to 125 KIAS and Temporary Pedal Stop

16 December 1998 10 March 1999

BHT-407-FM-1 FADEC Software Version 5.202 22 December 1998 17 December 2002

BHT-407-FM-1 Hover Performance Correction for Temporary Tail Rotor Pedal Stop

10 March 1999 17 December 2002

BHT-407-FM-1 VNE Increase to 130 KIAS Reissue3 June 1999

17 December 2002

BHT-407-FM-1 FADEC Direct Reversion to Manual System, ASB 407-99-31

4 June 1999 17 December 2002

BHT-407-FM-1 VNE Increase to 140 KIAS Revision 127 June 2000

17 December 2002

BHT-407-FM-1 (TR-9) Sustained Hover and Vertical Takeoff/Landing Operations with Tailwind

15 January 2002

BHT-407-FM-1 (TR-10) Incorporation of Oil Cooler Blower Inlet Ducts and Bearing Airflow Shields

Revision 125 July 2002

BHT-407-FM-1 (TR-11) FADEC Software Version 5.356 (Reversionary Governor)

16 November 2006

BHT-407-FM-1

16 November 2006

ROTORCRAFT FLIGHT MANUAL

TEMPORARY REVISION FOR

SUSTAINED HOVER AND VERTICAL TAKEOFF/LANDING OPERATIONS

WITH TAILWINDInsert these Temporary Revision pages next to like-numberpages in the basic Flight Manual.

DO NOT remove existing pages. DO NOT remove temporarypages until replacement pages are received or temporarypages are canceled.

NOTE

For tracking purposes, Temporary Revisions are now beingnumbered. This Temporary Revision is issued as TR-9,following previously issued Temporary Revisions.

THIS MANUAL SHALL BE IN THE HELICOPTER DURING ALL OPERATIONS

COPYRIGHTCOPYRIGHTBELL ® HELICOPTER TEXTRON INC.® HELICOPTER TEXTRON INC.AND BELLAND BELL HELICOPTER TEXTRON INC.HELICOPTER TEXTRON INC.A DIVISION OF TEXTRON CANADADIVISION OF TEXTRON CANADA LTD.

ALL RIGHTS RESERRIGHTS RESERVED

2002

COPYRIGHT NOTICE

A Subsidiary of Textron Inc.

POST OFFICE BOX 482 FORT WORTH, TEXAS 76101

Bell Helicopter

BHT-407-FM-1

09 FEBRUARY 1996TEMPORARY REVISION (TR-9) — 15 JANUARY 2002

BHT-407-FM-1

LOG OF REVISIONSTemporary (TR-9)...............................15 JAN 02

LOG OF PAGES

REVISION REVISIONNO. NO.PAGE PAGE

Title ........................................Temporary (TR-9)A .............................................Temporary (TR-9)C-D .........................................Temporary (TR-9)1-5 ..........................................Temporary (TR-9)1-7 ..........................................Temporary (TR-9)

TEMPORARY REVISION (TR-9) — 15 JAN 2002 A

CHANGED INFORMATION IN THIS TEMPORARY REVISION APPEARS IN BOLD TYPE.ALL OTHER INFORMATION IS OUTLINED.

NOTICE

BHT-407-FM-1

LOG OF TC APPROVED REVISIONS

Temporary (TR-9)...............................15 JAN 02

APPROVED DATE

CHIEF, FLIGHT TESTFORDIRECTOR — AIRCRAFT CERTIFICATIONTRANSPORT CANADA

TEMPORARY REVISION (TR-9) — 15 JAN 2002 C

CHIEF, FLIGHT TESTFORDIRECTOR — AIRCRAFT CERTIFICATION BRANCHDEPARTMENT OF TRANSPORT

BHT-407-FM-1

LOG OF FAA APPROVED REVISIONS

Temporary (TR-9) .............................. 18 JAN 02

D TEMPORARY REVISION (TR-9) — 15 JAN 2002

TC APPROVED BHT-407-FM-1

Maximum allowable airspeed for sidewardand rearward flight or crosswind hover is35 KTAS.

Sustained hover and vertical takeoff/landing operation (greater than one minute)with tailwind (relative winds within ± 90° oftail) greater than 5 knots is prohibited.

1-8. ALTITUDE

Maximum operating altitude is 20,000 feet Hp.

TEMPORARY REVISION (TR-9) — 15 JAN 2002 1-5


1-13-G. ENGINE OIL TEMPERATURE

Continuous operation 0 to 107°C

Maximum 107°C

CAUTION

IF HOVERING WITH A TAILWINDGREATER THAN 5 KNOTS ATO A T A B O V E 2 4 ° C ( 7 5 ° F ) ,C L O S E L Y M O N I T O R E N G I N EA N D T R A N S M I S S I O N O I LT E M P E R A T U R E S . I F E N G I N EO R T R A N S M I S S I O N O I LT E M P E R A T U R E S R I S EABNORMALLY, TURN INTO WIND,REDUCE POWER OR TRANSITIONTO FORWARD FLIGHT UNTILTEMPERATURE DECREASES.

NOTE

Positive temperature indication is whenthe second segment of the trend arc isilluminated.

1-14. TRANSMISSION

TEMPORARY REVISION (TR-9) — 15 JAN 2002 1-7

ROTORCRAFTFLIGHT MANUAL

TEMPORARY REVISION FOR THEINCORPORATION OF OIL COOLER BLOWER

INLET DUCTS AND BEARINGAIRFLOW SHIELDS

This Temporary Revision supersedes and replaces in its entirety, TemporaryRevision for Sustained Hover and Vertical Takeoff/Landing Operations withTailwind, TR-9 dated 15 January 2002, when Oil Cooler Blower Inlet Ducts andBearing Airflow Shields have been incorporated. DO NOT incorporate thisTemporary Revision into manual or remove previously issued TR-9, untilmodifications 407-799-057 (Inlet Ducts) and 407-799-055 (Bearing AirflowShields) or ASB 407-02-54 has been accomplished.

Helicopter S/N 53519 and subsequent will have these modificationsincorporated as basic configuration.

Insert these Temporary Revision pages next to like-number pages in the basicFlight Manual.

DO NOT remove existing pages. DO NOT remove temporary pages untilreplacement pages are received or temporary pages are canceled.

NOTE

For tracking purposes, Temporary Revisions are now being numbered. ThisTemporary Revision is issued as TR-10 Revision 1, following previouslyissued Temporary Revisions.


COPYRIGHTBELL ® HELICOPTER TEXTRON INC.® HELICOPTER TEXTRON INC.AND BELLAND BELL HELICOPTER TEXTRON INC.HELICOPTER TEXTRON INC.A DIVISION OF TEXTRON CANADADIVISION OF TEXTRON CANADA LTD.

ALL RIGHTS RESERRIGHTS RESERVED

2002

COPYRIGHT NOTICE



Bell Helicopter

BHT-407-FM-1

09 FEBRUARY 1996TEMPORARY REVISION (TR-10) — 15 FEBRUARY 2002

REVISION 1 —25 JULY 2002

BHT-407-FM-1

LOG OF REVISIONS

Temporary (TR-10).............................15 FEB 02 Revision ...................... 1 .....................25 JUL 02

LOG OF PAGES


Title ...........................................................1A ................................................................1C-D ............................................................1

1-7 ......................................... Temporary (TR-10)2-3 ......................................... Temporary (TR-10)2-5 — 2-6 .............................. Temporary (TR-10)

TEMPORARY REVISION (TR-10) — 15 FEB 2002 A

CHANGED INFORMATION IN THIS TEMPORARY REVISION APPEARS IN BOLD TYPE.ALL OTHER INFORMATION IS OUTLINED.

NOTICE

REVISION 1 — 25 JULY 2002

BHT-407-FM-1

APPROVED DATE


DATE


Temporary (TR-10).............................15 FEB 02 Revision ...................... 1 .....................25 JUL 02

TEMPORARY REVISION (TR-10) — 15 FEB 2002 C

CHIEF, FLIGHT TESTFORDIRECTOR — AIRCRAFT CERTIFICATION BRANCHDEPARTMENT OF TRANSPORT

REVISION 1 — 25 JULY 2002

BHT-407-FM-1


Temporary (TR-10) ............................ 22 FEB 02 Revision ..................... 1..................... 25 JUL 02

D TEMPORARY REVISION (TR-10) — 15 FEB 2002REVISION 1 — 25 JULY 2002




Maximum 107°C

If hovering with a tailwind greater than 10knots at OAT above 37.8°C (100°F),closely monitor engine oil temperature.The oil temperature may be reduced byeither turning into wind, reducing poweror transition to forward flight.

NOTE

Positive temperature indication is whenthe second segment of the trend arc isilluminated.

1-14. TRANSMISSION

TEMPORARY REVISION (TR-10) — 15 FEB 2002 1-7


NORMAL PROCEDURES

Section 2


2-1-B. HOT WEATHER OPERATIONS

CAUTION

IF HOVERING WITH A TAILWINDGREATER THAN 10 KNOTS ATO AT A B O V E 3 7 . 8 ° C ( 1 0 0 ° F ) ,C L O S E L Y M O N I T O R E N G I N EOIL TEMPERATURE. THE OILT E M P E R A T U R E M A Y B EREDUCED BY EITHER TURNINGINTO WIND, REDUCING POWEROR TRANSITION TO FORWARDFLIGHT.

2-2. FLIGHT PLANNING



f. Hydromechanical unit — Security and condition; evidence of leakage.

g. Hoses and tubing — Chafing, security, and condition.

h. Oil cooler blower inlet duct andscreen — Clear of obstructions,condition and security.

17. Engine cowl — Secured.

18. Generator cooling scoop — Clear of debris.

BHT-407-FM-1 TC APPROVED

2-6 TEMPORARY REVISION (TR-10) — 15 FEB 2002

f. Hydromechanical unit —Security and condition;evidence of leakage.

j. Rotor brake disc and caliper (ifinstalled) — Condition, securityo f a t t a c h m e n t a n d l e a k a g e.Ensure brake pads are retractedfrom brake disc.

j1. Oil cooler blower inlet duct andscreen — Clear of obstructions,condition and security.

k. Engine cowling — Secured.

l. Air induction cowling — Secured.


18. Generator cooling scoop —Clear of debris.

FADEC SOFTWAREVERSION 5.356

(REVERSIONARY GOVERNOR)

Insert these Temporary Revision pages opposite likenumbered pages in the basic Flight Manual after helicopteris configured with FADEC Software Version 5.356.

DO NOT remove existing pages from Flight Manual. DO NOTremove temporary pages until replacement pages arereceived or temporary pages are canceled.


TEMPORARY REVISIONFOR


BHT-407-FM-1


REISSUE — 17 DECEMBER 2002TEMPORARY REVISION (TR-11) — 16 NOVEMBER 2006

COPYRIGHT NOTICE

COPYRIGHT 2006

BELL ® HELICOPTER TEXTRON INC.

AND BELL HELICOPTER TEXTRON

CANADA LTD.

ALL RIGHTS RESERVED

BHT-407-FM-1

Temporary (TR-11)....0 ..................... 16 NOV 06

LOG OF REVISIONS

LOG OF PAGES


FLIGHT MANUAL

Title ................................ Temporary (TR-11)A/B.................................. Temporary (TR-11)C/D.................................. Temporary (TR-11)E/F .................................. Temporary (TR-11)1-3...................................Temporary (TR-11)1-15................................. Temporary (TR-11)2-9...................................Temporary (TR-11)

Temporary Revision (TR-11) — 16 NOV 2006———A/B

NOTICE

Changed information in this Temporary Revision appears in bold type.All other information is outlined.

APPROVED DATE

BHT-407-FM-1

TEMPORARY REVISION (TR-11) — 16 NOV 2006———C/D

Temporary (TR-11)....0 ..................... 16 NOV 06



BHT-407-FM-1

TEMPORARY REVISION (TR-11) — 16 NOV 2006———E/F

Temporary (TR-11) ...0 ..................... 16 NOV 06



TEMPORARY REVISION (TR-11) — 16 NOV 2006———1-3

Section 1LIMITATIONSTIONS

1

1-5. CONFIGURATION

1-5-A. REQUIRED EQUIPMENT

A functional flashlight is required for nightflights.

FADEC system software shall be version5.356.



Figure 1-3. Placards and decals (Sheet 3 of 3)

Location: Instrument panel.

FADEC SOFTWARE VERSION 5.356WITH DIRECT REVERSION TO

MANUAL INSTALLED, REFER TOFLIGHT MANUAL FOR OPERATION.

407FM_TR11_01



c. After 3.5 seconds; ENG OUT,FADEC DEGRADE, FADECFAULT, RESTART FAULT, andENGINE OVSPD lights illuminatewith activation of engine outaudio for 3 seconds.

d. Sequence repeats (secondtime).

e. ENG OUT light re-illuminationwith reactivation of engine outaudio after 3 seconds (thirdtime).

18. HORN MUTE button — Press tomute.

19. Caution lights — ENG OUT, XMSNOIL PRESS, RPM, HYDRAULICSYSTEM, GEN FAIL, L/FUEL BOOST,R/FUEL BOOST, L/FUEL XFR, andR/FUEL XFR will be illuminated.

BHT-407-FM-1

POST OFFICE BOX 482 •FORT WORTH, TEXAS 76101


BHT-407-FM-1

REISSUE — 17 DECEMBER 2002REVISION 7 — 30 JUL 2008



TYPE CERTIFICATE NO. ________

REGISTRATION NO. ______________________ SERIAL NO. ____________________

APPROVED BY DATE _____________________

DIRECTOR — AIRCRAFT CERTIFICATION BRANCHDEPARTMENT OF TRANSPORT

THE AVIATION REGULATORY AUTHORITY FOR THIS FLIGHT MANUAL IS THECANADIAN DEPARTMENT OF TRANSPORT, AIRCRAFT CERTIFICATION BRANCH.

U.S. REGISTERED HELICOPTERS ARE APPROVED BY THE FAA IN ACCORDANCE WITH THE PROVISIONS OF 14 CFR SECTION 21.29 ON

H-92

9 FEBRUARY 1996

23 FEBRUARY 1996

COPYRIGHT NOTICECOPYRIGHT 2008BELL ® HELICOPTER TEXTRON INC.AND BELL HELICOPTER TEXTRONCANADA LTD.

ALL RIGHTS RESERVED

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Temporary RevisionS 4

BHT-407-FM-1

The following Warning is not applicable to helicopters on which all kits and customizinginstallations have been qualified and approved by Bell Helicopter.

WARNING

THE HELICOPTER MAY CONTAIN INSTALLATIONS, PARTS, ORPROCESSES CERTIFIED BY PARTIES OTHER THAN BELLHELICOPTER TEXTRON. BELL HELICOPTER CAN NOTCONFIRM THAT SUCH INSTALLATIONS HAVE BEEN FULLYQUALIFIED OR CONFORMED TO BELL HELICOPTER DESIGNCRITERIA. AS A RESULT OF SUCH INSTALLATIONS, BELLHELICOPTER SUPPLIED DATA MAY NOT BE VALIDCONCERNING IN-FLIGHT HANDLING QUALITIES, WEIGHT ANDBALANCE, OR HELICOPTER PERFORMANCE. IF MULTIPLE STCKITS OR SIMILAR INSTALLATIONS ARE INCORPORATED,THERE MAY BE NOT VALID TEST DATA TO QUALIFY THEHELICOPTER AS MODIFIED BY THESE INSTALLATIONS. FORREVISED DATA, CONTACT THE OWNER OF THE INSTALLEDSTC OR THE SUPPLIER FOR THE APPLICABLE APPROVAL OFEACH INSTALLATION.

NP 17 DEC 2002

Additional copies of this publication may be obtained by contacting:Commercial Publication Distribution Center

Bell Helicopter Textron Inc.P. O. Box 482

Fort Worth, Texas 76101-0482

NOTICE PAGE

30 JUL 2008—Rev. 7———A/B

BHT-407-FM-1

LOG OF PAGES


NOTE

Revised text is indicated by a black vertical line. Insert latest revision pages; dispose of superseded pages.

Original ......................0 ......................09 FEB 96Revision.....................1 .....................08 MAR 96Revision.....................2 .....................09 MAY 96Revision.....................3 ...................... 30 JUL 96Revision.....................4 ..................... 04 NOV 96Revision.....................5 ......................24 JUN 97Revision.....................6 ...................... 03 JUL 98Revision.....................7 ......................04 SEP 98Revision.....................8 ..................... 17 APR 00

Reissue ..................... 0 ......................17 DEC 02Revision .................... 1 ..................... 10 MAR 04Revision .................... 2 ......................29 NOV 04Revision .................... 3 ......................26 APR 05Revision .................... 4 ...................... 29 JUN 05Revision .................... 5 ...................... 19 FEB 07Revision .................... 6 ...................... 20 JUN 07Revision .................... 7 .......................30 JUL 08

Cover............................................................ 0Title .............................................................. 7NP................................................................. 0A/B................................................................ 7C/D................................................................ 7E/F ................................................................ 7i – ii............................................................... 0iii/iv............................................................... 01-1 – 1-3 .......................................................51-4 – 1-8 .......................................................71-9 – 1-13 ..................................................... 01-14............................................................... 51-15............................................................... 71-16 – 1-17 ................................................... 01-18 – 1-19 ................................................... 71-20............................................................... 32-1/2-2 .......................................................... 62-3 – 2-7 .......................................................52-8 – 2-9 .......................................................72-10 – 2-11 ................................................... 5

2-12 – 2-15................................................... 62-16 .............................................................. 72-17/2-18...................................................... 03-1/3-2.......................................................... 73-3 ................................................................ 73-4 – 3-8....................................................... 03-9 ................................................................ 73-10 – 3-12................................................... 03-13 .............................................................. 53-14 – 3-17................................................... 03-18 – 3-19................................................... 53-20 .............................................................. 04-1/4-2.......................................................... 04-3 – 4-50..................................................... 04-51/4-52...................................................... 05-1 – 5-18..................................................... 05-19/5-20...................................................... 0A-1/A-2......................................................... 5A-3 – A-4...................................................... 5A-5/A-6......................................................... 5

LOG OF REVISIONS

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APPROVED DATE


30 JUL 2008—Rev. 7———C/D

BHT-407-FM-1

Original ..................... 0...................... 09 FEB 96Revision .................... 1..................... 08 MAR 96Revision .................... 2......................09 MAY 96Revision .................... 3.......................30 JUL 96Revision .................... 4......................04 NOV 96Revision .................... 5...................... 24 JUN 97Revision .................... 6.......................03 JUL 98Revision .................... 7...................... 04 SEP 98Revision .................... 8...................... 17 APR 00

Reissue......................0 ..................... 17 DEC 02Revision ....................1 .....................10 MAR 04Revision ....................2 ..................... 29 NOV 04Revision ....................3 ..................... 26 APR 05Revision ....................4 ......................29 JUN 05Revision ....................5 ......................19 FEB 07Revision ....................6 ......................20 JUN 07Revision ....................7 ...................... 30 JUL 08


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Original ..................... 0...................... 09 FEB 96Revision .................... 1..................... 08 MAR 96Revision .................... 2......................09 MAY 96Revision .................... 3.......................30 JUL 96Revision .................... 4......................04 NOV 96Revision .................... 5...................... 24 JUN 97Revision .................... 6......................18 MAY 99Revision .................... 7......................18 MAY 99Revision .................... 8......................05 MAY 00

Reissue......................0 ......................24 SEP 03Revision ....................1 .....................11 MAR 04Revision ....................2 ..................... 07 DEC 04Revision ....................3 ..................... 04 MAY 05Revision ....................4 ......................30 JUN 05Revision ....................5 .....................08 MAR 07Revision ....................6 ..................... 31 AUG 07Revision ....................7 ..................... 18 AUG 08

30 JUL 2008—Rev. 7———E/F

BHT-407-FM-1

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BHT-407-FM-1

GENERAL INFORMATION

ORGANIZATION

This Rotorcraft Flight Manual is divided intofive sections and an appendix as follows:

Sections 1 through 4 contain TransportCanada (TC) approved data necessary tooperate basic helicopter in a safe andefficient manner.

Section 5 contains weight and balance datanecessary for flight planning.

Appendix A contains a list of approvedsupplements for optional equipment, whichshall be used in conjunction with basicFlight Manual when respective optionalequipment kits are installed.

Manufacturer's Data manual (BHT-407-MD-1)contains information to be used inconjunction with Flight Manual.Manufacturer's data manual is divided intofour sections:

TERMINOLOGY

WARNINGS, CAUTIONS, AND NOTES

Warnings, cautions, and notes are usedthroughout this manual to emphasizeimportant and critical instructions asfollows:

WARNING

AN OPERATING PROCEDURE,PRACTICE, ETC., WHICH, IF NOTCORRECTLY FOLLOWED, COULDRESULT IN PERSONAL INJURY ORLOSS OF LIFE.

CAUTION

AN OPERATING PROCEDURE,PRACTICE, ETC., WHICH IF NOTSTRICTLY OBSERVED, COULDRESULT IN DAMAGE TO ORDESTRUCTION OF EQUIPMENT.

NOTE

An operating procedure, condition,etc., which is essential to highlight.

USE OF PROCEDURAL WORDS

Concept of procedural word usage andintended meaning which has been adheredto in preparing this manual is as follows:

SHALL has been used only whenapplication of a procedure is mandatory.

SHOULD has been used only whenapplication of a procedure is recommended.

Section 1 — LIMITATIONSSection 2 — NORMAL PROCEDURESSection 3 — EMERGENCY AND

MALFUNCTION PROCEDURES

Section 4 — PERFORMANCESection 5 — WEIGHT AND BALANCEAppendix A — OPTIONAL EQUIPMENT

SUPPLEMENTS

Section 1 — SYSTEMS DESCRIPTION

Section 2 — HANDLING AND SERVICING

Section 3 — CONVERSION CHARTS AND TABLES

Section 4 — EXPANDED PERFORMANCE

17 DEC 2002 i

BHT-407-FM-1

MAY and NEED NOT have been used onlywhen application of a procedure is optional.

WILL has been used only to indicatefuturity, never to indicate a mandatoryprocedure.

ABBREVIATIONS, ACRONYMS ANDPLACARDING

Abbreviations, acronyms and placardingused throughout this manual are defined asfollows:

ADF — Automatic direction finder

AIRCOND

— Air conditioner

A/F — Airframe

ALT — Altimeter

ANTICOLL LT

— Anticollision light

ATT — Attitude

AUTO — Automatic

AUX — Auxiliary

BATT — Battery

BIT — Built in test

BL — Buttock line

BLO — Blower

BRT — Bright

°C — Degrees Celsius

CAUT — Caution

CAUT LT — Caution lights

CG — Center of gravity

CKPT — Cockpit

CM — Centimeter (s)

COMM — Communication

CONT — Control

dBA — Decibel, “A” type filter

DC — Direct current

DG — Directional gyro

DOT — Department of Transport

ECS — Environmental control system

ECU — Engine control unit

ELT — Emergency locator transmitter

ENCDG — Encoding

ENG — Engine

ENGANTI ICE

— Engine anti icing

°F — Degrees Fahrenheit

FADEC — Full authority digital engine control

FS — Fuselage station

FT or ft — Foot, feet

FWD — Forward

GEN — Generator

GOV — Governor

GPS — Global positioning system

GPU — Ground power unit

GW — Gross weight

HD — Density altitude

HG — Inches of mercury

HMU — Hydromechanical unit

Hp — Pressure altitude

HYD — Hydraulic

HV — Height-velocity

ICAO — International Civil Aviation Organization

ICS — Intercommunication system

IFL — Inflate

ii 17 DEC 2002

BHT-407-FM-1

IGE — In ground effect

IGNTR — Ignitor

IN — Inch(es)

INSTRCHK

— Instrument check

INSTR LT — Instrument light

KCAS — Knots calibrated airspeed

KG or kg — Kilogram(s)

KIAS — Knots indicated airspeed

KTAS — Knots true airspeed

L — Liter(s)

LB(S) or lb(s)

— Pound(s)

LDG LTS — Landing lights

L/FUEL — Left fuel

LT — Light

MAN — Manual

MCP — Maximum continuous power

MD — Manufacturer's Data

MGT — Measured gas temperature

MM or mm

— Millimeter(s)

NAV — Navigation

NG — Gas producer RPM

NP — Power turbine RPM

NR — Rotor RPM

OAT — Outside air temperature

OBS — Omni bearing selector

OGE — Out of ground effect

OVSPD — Overspeed

PARTSEP

— Particle separator

PASS — Passenger(s)

PMA — Permanent Magnetic Alternator

POS LT — Position light

PRESS — Pressure

PSI — Pounds per square inch

PTT — Press to Test

PWR — Power

QTY — Quantity

R/FUEL — Right fuel

RECP — Receptacle

RLY — Relay

RPM — Revolutions per minute

RTR — Rotor

s/w Ver Soft ware version

SEL — Sound exposure level

SHP — Shaft horsepower

SL — Sea level

SPKR — Speaker

Sq — Square

SYS — System

T/R — Tail rotor

TCA — Transport Canada Aviation

TEMP — Temperature

TRQ — Torque

VFR — Visual flight rules

VHF — Very high frequency

VNE — Never exceed velocity

VOR — VHF omnidirectional range

WL — Water line

WARN — Warning

XFR — Transfer

XMSN — Transmission

XPDR — Transponder

17 DEC 2002 iii/iv


19 FEB 2007—Rev. 5———1-1

Section 1LIMITATIONS

1TABLE OF CONTENTS

Paragraph PageSubject Number Number

Introduction ............................................................................................ 1-1 ........... 1-3Basis of Certification ............................................................................. 1-2 ........... 1-3Types of Operation ................................................................................ 1-3 ........... 1-3

Passengers......................................................................................... 1-3-A ....... 1-3Cargo................................................................................................... 1-3-B ....... 1-3

Flight Crew ............................................................................................. 1-4 ........... 1-3Configuration ......................................................................................... 1-5 ........... 1-3

Required Equipment.......................................................................... 1-5-A ....... 1-3 Optional Equipment.......................................................................... 1-5-B ....... 1-4 Doors Removed ................................................................................ 1-5-C ....... 1-4

Weight and Center of Gravity ............................................................... 1-6 ........... 1-4Weight ................................................................................................. 1-6-A ....... 1-4Center of Gravity................................................................................ 1-6-B ....... 1-4

Airspeed.................................................................................................. 1-7 ........... 1-4Altitude.................................................................................................... 1-8 ........... 1-5Maneuvering........................................................................................... 1-9 ........... 1-5

Prohibited Maneuvers ....................................................................... 1-9-A ....... 1-5Climb And Descent ............................................................................ 1-9-B ....... 1-5Slope Landing .................................................................................... 1-9-C ....... 1-5

Not Used ................................................................................................. 1-10 ......... 1-5Ambient Temperatures.......................................................................... 1-11 ......... 1-5Electrical ................................................................................................. 1-12 ......... 1-5

Generator............................................................................................ 1-12-A ..... 1-5Starter ................................................................................................. 1-12-B ..... 1-5

Power Plant ............................................................................................ 1-13 ......... 1-5Gas Producer RPM (NG) .................................................................... 1-13-A ..... 1-6Power Turbine RPM (NP) ................................................................... 1-13-B ..... 1-6Measured Gas Temperature (MGT) .................................................. 1-13-C ..... 1-6Engine Torque.................................................................................... 1-13-D ..... 1-6Fuel Pressure ..................................................................................... 1-13-E ..... 1-6Engine Oil Pressure........................................................................... 1-13-F...... 1-6Engine Oil Temperature .................................................................... 1-13-G..... 1-7

Transmission.......................................................................................... 1-14 ......... 1-7Transmission Oil Pressure ............................................................... 1-14-A ..... 1-7Transmission Oil Temperature ......................................................... 1-14-B ..... 1-7

Rotor ....................................................................................................... 1-15 ......... 1-7


1-2———Rev. 5—19 FEB 2007

TABLE OF CONTENTS (CONT)


Rotor RPM — Power ON ................................................................... 1-15-A..... 1-7Rotor RPM — Power OFF.................................................................. 1-15-B..... 1-7

Hydraulic ................................................................................................ 1-16......... 1-7Fuel and Oil ............................................................................................ 1-17......... 1-7

Fuel ..................................................................................................... 1-17-A..... 1-7Oil ........................................................................................................ 1-17-B..... 1-8

Rotor Brake ............................................................................................ 1-18......... 1-8Not Used................................................................................................. 1-19......... 1-8Instrument Markings and Placards ...................................................... 1-20......... 1-8

LIST OF FIGURES

Figure PageSubject Number NumberGross Weight Longitudinal Center of Gravity Limits ......................... 1-1........... 1-9Gross Weight Lateral Center of Gravity Limits................................... 1-2........... 1-11Placards and Decals.............................................................................. 1-3........... 1-13Ambient Air Temperature Limitations ................................................. 1-4........... 1-16Instrument Markings ............................................................................. 1-5........... 1-17


19 FEB 2007—Rev. 5———1-3


11-1. INTRODUCTION

Compliance with Limitations section isrequired by appropriate operating rules.Anytime an operating limitation is exceeded,an appropr iate entry shal l be made inhelicopter logbook. Entry shall state whichlimit was exceeded, duration of time, extremevalue attained, and any additional informationessential in determining maintenance actionrequired.

In ten t iona l use o f t rans ient l imi ts isprohibited.

Torque events shall be recorded. A torqueevent is defined as a takeoff or lift, internal orexternal load (BHT-407-MD-1).

Landings shall be recorded. Run-on landingsshall be recorded separately.

A run-on landing is defined as one wherethere is fo rward ground t rave l o f thehelicopter greater than 3 feet with the weighton the skids.

1-2. BASIS OF CERTIFICATION

This helicopter is certified under FARs Parts27 and 36, Appendix J. Additionally, it isapproved under Canadian AirworthinessManual Chapters 516 (ICAO Chapter 11) and527, Sections 1093 (b) (1) (ii) and (iii), 1301-1,1557 (c) (3), 1581 (e) and 1583 (h).

1-3. TYPES OF OPERATION

1-3-A. PASSENGERS

Basic configured helicopter is approved forseven place seating and is certified for land

operation under day or night VFR non-icingconditions.

1-3-B. CARGO

The maximum allowable cabin deck loadingfor cargo is 75 pounds per square foot (3.7 kgper 100 cm2). The maximum al lowablebaggage compartment deck loading is 86pounds per square foot (4.2 kg per 100 cm2)with a maximum allowable weight of 250pounds (113.4 kg). Refer to BHT-407-MD-1 forcargo restraint and tie-down locations.

Cargo must be properly secured by tie-downdevices to prevent the load from shiftingunder ant ic ipated f l igh t and g roundoperations. If the mission requires bothpassengers and cargo to be transportedtogether, the cargo must be loaded andsecured so tha t i t does no t obs t ructpassenger access to exits.

1-4. FLIGHT CREW

Minimum flight crew consists of one pilot whoshall operate helicopter from right crew seat.

Left crew seat may be used for an additionalpi lot when approved dual controls areinstalled.

1-5. CONFIGURATION


A functional flashlight is required for nightflights.

FADEC system software shall be version5.202.

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1-4———Rev. 7—30 JUL 2008

1-5-B. OPTIONAL EQUIPMENT

The snow deflector kit (BHT-407-FMS-4) shallbe ins ta l led when conduct ing f l ightoperations in falling and/or blowing snow.

Refer to appropr ia te f l ight manualsupplement (s ) (FMS) for addi t iona llimitations, procedures, and performance datafor optional equipment.

1-5-C. DOORS REMOVED

NOTE

Indicated altitude may be up to 100feet lower than actual altitude withcrew door(s) removed.

Flight with any combination of doors removedis approved. With litter door removed, leftpassenger door shall be removed. Refer toAirspeed limitations.

With door(s) removed, determine weightchange and adjust ballast if necessary. Referto Section 5.

NOTE

Al l unsecured i tems sha l l beremoved from cabin when any dooris removed.

1-6. WEIGHT AND CENTER OFGRAVITY

1-6-A. WEIGHT

Maximum approved internal GW for takeoffand landing is 5000 pounds (2268 kg).

Minimum GW for f l ight is 2650 pounds(1202 kg).

Minimum weight at fuselage station 65.0 is170 pounds (77.1 kg).

CAUTION

LOADS THAT RESULT IN GWABOVE 5000 POUNDS (2268 KG)SHALL BE CARRIED ON THECARGO HOOK AND MUST BEJETTISONABLE.

Maximum approved GW for f l ight wi thjettisonable external load is 6000 pounds(2722 kg).

1-6-B. CENTER OF GRAVITY

The pilot is responsible for determiningweight and balance to ensure gross weightand center of gravity will remain within limitsthroughout each flight. Refer to Section 5 forloading tables and instructions.

NOTE

Ballast as required to maintain mostforward or most aft CG within GWflight limits (Figure 1-1). For standardpassenger and fue l load ings ,applicable Weight Empty Center ofGravity Chart in BHT-407-MM-1 maybe used to determine requiredballast.

For longitudinal CG limits, refer to GrossWeight Longitudinal Center of Gravity Limitschart (Figure 1-1).

For lateral CG limits, refer to Gross WeightLateral Center of Gravity Limits (Figure 1-2).

1-7. AIRSPEED

Basic VNE is 140 KIAS, sea level to 3000 feetHD. Decrease VNE for ambient conditions inaccordance with AIRSPEED LIMITATIONSPlacards and Decals (Figure 1-3).

VNE at 93.5 to 100% TORQUE (takeoff power)is 100 KIAS, not to exceed placarded VNE.


30 JUL 2008—Rev. 7———1-5

VNE is 100 KIAS or placarded VNE, whicheveris less, when takeoff loading is in shaded areaof the Gross Weight Lateral Center of GravityLimits (Figure 1-2).

VNE is 100 KIAS with any door(s) removed, notto exceed placarded VNE.

VNE is 100 KIAS or placarded VNE, whicheveris less for steady state autorotation.

Maximum allowable airspeed for sidewardand rearward flight or crosswind hover is 35KTAS.

1-8. ALTITUDE

Maximum operating altitude is 20,000 feet HDor 20,000 feet HP, whichever is lower.

1-9. MANEUVERING

1-9-A. PROHIBITED MANEUVERS

Aerobatic maneuvers are prohibited.

1-9-B. CLIMB AND DESCENT

Maximum rate of climb is 2000 feet perminute.

1-9-C. SLOPE LANDING

CAUTION

SLOPE LANDINGS HAVE BEENDEMONSTRATED TO THE SLOPELANDING L IMITS. OTHERCONDITIONS INCLUDING, BUT NOTLIMITED TO, WIND DIRECTION ANDVELOCITY, CENTER OF GRAVITY,AND THE CONDIT ION OF THESLOPE (LOOSE ROCK, SOFT MUD,SNOW, WET GRASS, ETC.) MAYLIMIT MAXIMUM SLOPE TO A VALUELESS THAN THE PUBLISHEDLIMITS.

Slope landings are limited to 10° side slopes,10° nose up slope or 5° nose down slope.

1-10. NOT USED

1-11. AMBIENT TEMPERATURES

Maximum sea level ambient air temperaturefor operation is 51.7°C (125°F) and decreaseswith HP at standard lapse rate of 2°C (3.6°F)per 1000 fee t . Refer to Ambien t A i rTemperature Limitations chart (Figure 1-4).

Min imum ambient a i r temperature foroperation at all altitudes is -40°C (-40°F).

ENG ANTI ICE shall be ON in visible moisturewhen OAT is below 5°C (40°F).

1-12. ELECTRICAL

1-12-A. GENERATOR

1-12-B. STARTER

Continuous operation,up to 10,000 feet Hp

0 to 180 amps

Maximum continuous up to 10,000 feet Hp

180 amps

Continuous operation,above 10,000 feet Hp

0 to 170 amps

Maximum continuous above 10,000 feet Hp

170 amps

Transient, 2 minutes 180 to 300 amps

Transient, 5 seconds 300 to 400 amps

External Power Start Battery Start

40 seconds ON 60 seconds ON

30 seconds OFF 60 seconds OFF


30 seconds OFF 60 seconds OFF


30 minutes OFF 30 minutes OFF

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1-6———Rev. 7—30 JUL 2008

NOTE

28 VDC GPU for starting shall belimited to 500 amps.

1-13. POWER PLANT

Rolls-Royce model 250-C47B.

NOTE

Inten t iona l use o f any powertransient is prohibited.

1-13-A. GAS PRODUCER RPM (NG)

NOTE

FADEC will limit NG in accordancewith GAS PRODUCER RPM (NG)LIMIT placard. NR decay will result ifpower demand exceeds placard limit.

Maximum cont inuous NG is l im i ted inaccordance with GAS PRODUCER RPM (NG)LIMIT placard (Figure 1-3) when operatingabove 10,000 feet HP and with OAT below-30°C (-22°F).

1-13-B. POWER TURBINE RPM (NP)

NOTE

ENGINE OVSPD warning light willilluminate when NP versus TORQUEis between 102.4% NP at 100%TORQUE and 108.6% NP at 0%TORQUE.

When operating in MANUAL mode NPshould be maintained between 95and 100%.

1-13-C. MEASURED GAS TEMPERATURE(MGT)

GAUGE P/N 407-375-001-101/-103

NOTE

Either MGT gauge may be installed.

GAUGE P/N 407-375-001-105 AND SUB

Continuous operation 63 to 105%

Maximum continuous operation

105%

Transient, 10 seconds 105.1 to 106%

Avoid continuousoperations

68.4 to 87.1%

Minimum 99%


Maximum continuous 100%

Maximum transient, 15 seconds

102.1 to 107% NP


Maximum continuous 727°C

Takeoff, 5 minutes 727 to 779°C

Maximum for takeoff 779°C

Transient, 12 seconds 780 to 826°C

Maximum starting, do not exceed 10 seconds above 826°C or 1 second at 927°C.

927°C


Maximum continuous 727°C

Takeoff, 5 minutes 727 to 779°C

Maximum for takeoff 779°C

Transient, 12 seconds 780 to 905°C

Maximum starting, do not exceed 10 seconds above 843°C or 1 second at 927°C.

927°C


30 JUL 2008—Rev. 7———1-7

1-13-D. ENGINE TORQUE

NOTE

Use of takeoff power is limited to 100KIAS, not to exceed placarded VNE.

1-13-E. FUEL PRESSURE

1-13-F. ENGINE OIL PRESSURE

NOTE

When 130 PSI is exceeded duringstart, operate engine at idle until oilpressure drops below 130 PSI.


NOTE

Positive temperature indication iswhen the second segment of thetrend arc is illuminated.

1-14. TRANSMISSION

1-14-A. TRANSMISSION OIL PRESSURE

1-14-B. TRANSMISSION OILTEMPERATURE

1-15. ROTOR

1-15-A. ROTOR RPM — POWER ON

NOTE

When operating in MANUAL modeNR should be maintained between95% and 100%.

1-15-B. ROTOR RPM — POWER OFF

CAUTION

FOR AUTOROTATIVE TRAINING,MAINTAIN STEADY STATE N RABOVE 90%.

1-16. HYDRAULIC

Hydraulic fluid MIL-PRF-5606 (NATO H-515)may be used at all ambient temperatures.

Continuous operation 0 to 93.5%

Maximum continuous 93.5%

Takeoff, 5 minute 93.5 to 100%

Transient, 5 seconds 105%

Minimum 8 PSI

Continuous operation 8 to 25 PSI

Maximum 25 PSI

Minimum below 79% NG 50 PSI

Minimum from 79 to 94% NG

90 PSI

Minimum above 94% NG 115 PSI

Maximum 130 PSI

Maximum cold starts only 200 PSI


Maximum 107°C

Minimum 30 PSI

Continuous operation 40 to 70 PSI

Maximum 70 PSI


Maximum 110°C


Maximum continuous 100%

Minimum 85%


Maximum 107%

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1-8———Rev. 7—30 JUL 2008

1-17. FUEL AND OIL

1-17-A. FUEL

Fuel conforming to following specificationsmay be used at all ambient temperatures:

ASTM-D-6615, Jet B

MIL-DTL-5624, Grade JP-4 (NATO F-40)

Fuels conforming to following specificationsare limited to ambient temperatures of -32°C(-25°F) and above:

ASTM-D-1655, Jet A or A-1

MIL-DTL-5624, Grade JP-5 (NATO F-44)

MIL-DTL-83133, Grade JP-8 (NATO F-34).

For operations below -32°C (-25°F), refer toRolls-Royce Operation and MaintenanceManual for cold weather fuel and blendinginstructions.

1-17-B. OIL

1-17-B-1. OIL — ENGINE

Oil conforming to MIL-PRF-7808 (NATOO-148), DOD-PRF-85734 or MIL-PRF-23699(NATO O-156) is l im i ted to ambienttemperatures above -40°C (-40°F).

NOTE

Refer to Rolls-Royce Operation andMaintenance Manual andBHT-407-MD-1 manual for approvedoils and mixing of oils of differentbrands, types, and manufacturers.

1-17-B-2. OIL — TRANSMISSION AND TAILROTOR GEARBOX

NOTE

It is recommended DOD-PRF-85734oil be used in transmission and tailrotor gearbox to maximum extentallowed by temperature limitations.

Oil conforming to DOD-PRF-85734 is limitedto ambient temperatures above -40°C (-40°F).

Oil conforming to MIL-PRF-7808 (NATO O-148)is limited to ambient temperatures below-18°C (0°F).

1-18. ROTOR BRAKE

Rotor brake (if installed) application is limitedto ground operation after engine has beenshut down and NR has decreased to 40% orlower.

For emergency stops, apply rotor brake anytime after engine is shut down.

Engine starts with rotor brake engaged areprohibited.

1-19. NOT USED

1-20. INSTRUMENT MARKINGSAND PLACARDS

Refer to Figure 1-3 for Placards and Decals.Refer to Figure 1-5 for Instrument Markings.

Illustrations shown in Figure 1-5 are artistrepresentations and may or may not depictactual approved instruments due to printinglimitations. Instrument operating ranges andlimits shall agree with those presented in thissection.


17 DEC 2002 1-9

Figure 1-1. Gross weight longitudinal center of gravity limits (Sheet 1 of 2)

118 119 120 121 122 123 124 125 126 127 128 129 130

FUSELAGE STATION - INCHES

2400

2600

2800

3000

3200

3400

3600

3800

4000

4200

4400

4600

4800

5000

5200

5400

5600

5800

6000

6200

GR

OS

S W

EIG

HT

- P

OU

ND

S

6000

5000

4500

2800

2650

120.5 127.6

119.5

119.0

128.0 129.0

EXTERNAL LOAD ONLY

LONGITUDINAL C.G.

M407_FM-1__FIG_1-1_(1_OF_2).WMF


1-10 17 DEC 2002


3000 3025 3050 3075 3100 3125 3150 3175 3200 3225 3250 3275 3300

FUSELAGE STATION - MILLIMETERS

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

2200

2300

2400

2500

2600

2700

2800

GR

OS

S W

EIG

HT

- K

ILO

GR

AM

S

LONGITUDINAL C.G.

2268

EXTERNAL LOAD ONLY

3241 3061

2722

3035

2041

1270

1202 3023

3251 3277

M407_FM-1__FIG_1-1_(2_OF_2).WMF


17 DEC 2002 1-11

Figure 1-2. Gross weight lateral center of gravity limits (Sheet 1 of 2)

-5 -4 -3 -2 -1 0 1 2 3 4 5

BUTTOCK LINE - INCHES

2400

2600

2800

3000

3200

3400

3600

3800

4000

4200

4400

4600

4800

5000

5200

5400

5600

5800

6000

6200

GR

OS

S W

EIG

HT

- P

OU

ND

S

5000

-1.5

LATERAL C.G.

EXTERNAL LOAD ONLY

MAX AIRSPEED 100 KIAS

-0.9 1.4

2.0

-4.0 -2.5 3.0 4.0

3500

2650

3500

6000

EXTERNALLOAD ONLY

MAX AIRSPEED100 KIAS

EXTERNALLOAD ONLY


M407_FM-1__FIG_1-2_(1_OF_2).WMF


1-12 17 DEC 2002


-125 -100 -75 -50 -25 0 25 50 75 100 125

BUTTOCK LINE - MILLIMETERS

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

2200

2300

2400

2500

2600

2700

2800

GR

OS

S W

EIG

HT

- K

ILO

GR

AM

S

LATERAL C.G.

EXTERNAL LOAD ONLY


EXTERNALLOAD ONLY


EXTERNALLOAD ONLY


-102 -64 76 102

1588 1588

1202

-39 52

2268

-23 36

2722

M407_FM-1__FIG_1-2_(2_OF_2).WMF


17 DEC 2002 1-13


Airspeed limits shown are valid only for corresponding altitudes and temperatures. Hatched areas indicate conditions which exceed approved temperature or density altitude limitations.

Location: Between Pilot and Copilot seats

0 2 4 6 8 10 12 14 16 18 2052 13745 139 132 12540 140 133 126 11935 140 135 128 120 11330 140 137 129 122 115 10825 140 138 131 124 116 109 102 9520 140 140 133 125 118 111 103 96 890 140 140 140 132 125 117 110 103 95 88

-25 140 140 140 135 130 125 119 111 104 97 89-40 137 133 128 123 118 114 110 105 101 97 93

PRESSURE ALTITUDE FT x 1000407 AIRSPEED LIMITATIONS - KIAS

OAT °C

MAXIMUM AUTOROTATION VNE 100 KIAS

EMERGENCY PEDALSTOP RELEASE

PULL ONLY MAINT. RESET

REQUIRED

Location: Forward of Overhead Console

M407_FM-1__FIG_1-3_(VNE).WMF


1-14 Rev. 5 19 FEB 2007


FUELFUEL SYSTEM USABLE CAPACITY

BASIC AIRCRAFT 127 U.S. GALLONS - 483 LITERSWITH 407-706-011 AUX KIT 147 U.S. GALLONS = 559 LITERS

SEE FLIGHT MANUAL FOR APPROVED FUELS

Location: Above fuel filler cap.

THIS HELICOPTER MUST BE OPERATED INCOMPLIANCE WITH THE OPERATING LIMITATIONS

SPECIFIED IN THE APPROVED FLIGHT MANUAL

Location: Bottom and centered on instrument panel.

CARGO MUST BE SECUREDIN ACCORDANCE WITHFLIGHT MANUAL INSTR

Location: Inside of baggage door.

407-FM-1-3-2

DO NOT APPLY ROTOR BRAKEABOVE 40% RPM

Location: Near rotor brake (if installed).

Location: Instrument panel.

AVOID CONT OPS 68.4% TO 87.1% NP


30 JUL 2008 Rev. 7 1-15

Figure 1-3. Placards and Decals (Sheet 3 of 3)

FADEC SOFTWARE VERSION 5.202

WITH DIRECT REVERSION TO

MANUAL INSTALLED. REFER TO

FLIGHT MANUAL FOR OPERATION

Location: Instrument panel

MAX ALLOWABLE WEIGHT 250 LBS.

MAX ALLOWABLE WEIGHT PER SQ. FT. 86 LBS.

Location: Inside of baggage door

FUEL CAPACITY

BASIC 869 LBS

WITH AUX 1005 LBS

(JET A AT 15ºC)

Location: Instrument panel

GAS PRODUCER RPM (NG) LIMIT

WHEN ABOVE 10,000 FT HP

MAXIMUM Ng % RPM WITH OAT IS AS FOLLOWS

Location: Above pilot windshield

OAT

ºC

MAX

Ng %99.0

-40

99.2

-39

99.4

-38

99.6

-37

99.8

-36

100.0

-35

100.2

-34

100.4

-33

100.6

-32

100.8

-31

101.1

-30

Location: Instrument panel and passenger compartment

407_FM_1_0002

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1-16 17 DEC 2002

Figure 1-4. Ambient air temperature limitations

-40 -30 -20 -10 0 10 20 30 40 50 60

AMBIENT AIR TEMPERATURE - °C

0

2,000

4,000

6,000

8,000

10,000

12,000

14,000

16,000

18,000

20,000

PR

ES

SU

RE

AL

TIT

UD

E

- F

T

-40 -20 0 20 40 60 80 100 120 140

AMBIENT AIR TEMPERATURE - °F M

INIM

UM

OA

T F

LIG

HT

LIM

IT

MA

XIM

UM

OA

T

EXAMPLE: AT HP = 7000 FTMAX OAT = 37.8°C (100°F)MIN OAT = - 40°C (- 40°F)

FLIGH

T LIMIT

MAXIMUM HD FLIGHT LIMIT

M407_FM-1__FIG_1-4.EMF


17 DEC 2002 1-17

Figure 1-5. Instrument markings (Sheet 1 of 4)

ENGINE OIL PRESSURE

50 PSI

50 to 90 PSI

90 to 115 PSI

115 to 130 PSI

130 PSI

200 PSI

ENGINE OIL TEMPERATURE

0 to 107°C

107°C

Minimum

Operation below 79% NG RPM

Continuous operation below 94% NG RPM

Maximum for continuous operation

Maximum for cold start

Continuous operation

Maximum


TRANSMISSION OIL PRESSURE

30 PSI

40 to 70 PSI

70 PSI

TRANSMISSION OIL TEMPERATURE

15 to 110°C

110°C

Minimum


Maximum


Maximum

NG (GAS PRODUCER RPM)

63 to 105%

105%

106%


Maximum continuous operation


M407_FM-1__FIG_1-5_(1_OF_4).EPS


1-18 Rev. 7 30 JUL 2008

Figure 1-5. Instrument Markings (Sheet 2 of 4)

TRQ (TORQUE)

MGT (MEASURED GAS TEMPERATURE)


5 minute takeoff range

Maximum for takeoff

Beginning of 10 seconds

range for starting

Maximum for start and shutdown

(1 second maximum)

407_FM_1_0003


Maximum for autorotation

Maximum

407-375-001-105 AND SUB.

Either gauge may be installed*

*

AIRSPEED

0 to 93.5%

93.5 to 100%

100%

0 to 140 Knots

100 Knots

140 Knots

100 to 727°C

727 to 779°C

779°C

826°C

or

843°C

927°C


5 minute takeoff range

Maximum


30 JUL 2008 Rev. 7 1-19

Figure 1-5. Instrument Markings (Sheet 3 of 4)

99%

99 to 100%

100%

Minimum


Maximum continuous

NP (POWER TURBINE RPM)

407_FM_1_0004

FUEL QUANTITY

0 LBS

195 LBS

869 LBS

1005 LBS

(Jet A 6.8 lbs/gal)

All tanks empty (zero useable)

Forward tank empty

Forward and aft tanks full

Forward, aft and auxiliary

tanks full

85%

85 to 107%

107%

Minimum (power off)

Continuous operation (power off)

Maximum (power off)

NR (ROTOR RPM)

407-375-008-101/-103


1-20 Rev. 3 26 APR 2005

Figure 1-5. Instrument markings (Sheet 4 of 4)

407FM_1_0001

DC LOAD

DC LOAD

FUEL PRESSURE

FUEL PRESSURE

* Either gauge may be installed.

* P/N 407-075-024-101

* P/N 407-075-024-103, 407-375-007-105 or 407-375-007-107

170 Amps

0 to 180 Amps

180 Amps

180 Amps

170 Amps

300 Amps

400 Amps

8 PSI

8 PSI

8 to 25 PSI

8 to 25 PSI

25 PSI

25 PSI

Maximum continuous above 10,000 FT Hp


Maximum

Maximum


Maximum transient, 2 minutes




Maximum

Maximum

Minimum

Minimum

VERTICAL SPEED INDICATOR

2,000 Feet per minute up Maximum


Section 2NORMAL PROCEDURES

2TABLE OF CONTENTS


Introduction ............................................................................................ 2-1 ........... 2-3 Cold Weather Operations................................................................. 2-1-A ....... 2-3 Hot Weather Operations................................................................... 2-1-B ....... 2-3

Flight Planning ....................................................................................... 2-2 ........... 2-3Preflight Check....................................................................................... 2-3 ........... 2-3

Before Exterior Check ...................................................................... 2-3-A ....... 2-4 Exterior Check................................................................................... 2-3-B ....... 2-4

Interior and Prestart Check................................................................... 2-4 ........... 2-7Engine Start............................................................................................ 2-5 ........... 2-9

Dry Motoring Run............................................................................... 2-5-A ....... 2-10Alternate Engine Start ....................................................................... 2-5-B ....... 2-10

Systems Check ...................................................................................... 2-6 ........... 2-11Preliminary Hydraulic Systems Check ............................................ 2-6-A ....... 2-11FADEC Manual Check ....................................................................... 2-6-B ....... 2-12Engine Runup..................................................................................... 2-6-C ....... 2-12Hydraulic Systems Check ................................................................. 2-6-D ....... 2-12

Before Takeoff........................................................................................ 2-7 ........... 2-13Takeoff .................................................................................................... 2-8 ........... 2-13In-Flight Operations............................................................................... 2-9 ........... 2-14Descent and Landing............................................................................. 2-10 ......... 2-14Engine Shutdown................................................................................... 2-11 ......... 2-15Postflight Check..................................................................................... 2-12 ......... 2-16

LIST OF FIGURES

Figure PageSubject Number NumberPreflight Check Sequence..................................................................... 2-1 ........... 2-17

20 JUN 2007—Rev. 6———2-1/2-2


19 FEB 2007—Rev. 5———2-3


22-1. INTRODUCTION

This section contains instructions andprocedures for operating helicopter fromplanning s tage , through actua l f l ightconditions, to securing helicopter afterlanding.

Normal and standard conditions are assumedin these procedures. Pertinent data in othersections is referenced when applicable.

Instructions and procedures contained hereinare written for purpose of standardization andare not applicable to all situations.

2-1-A. COLD WEATHER OPERATIONS

Battery starts have been demonstrated to-29°C (-20°F) with standard 17 amp-hourbattery and -35°C (-31°F) with optional 28amp-hour battery.

During engine start in cold temperatures,initial engine oil pressure of 200 PSI andpressure excursions down to 50 PSI duringwarm up are normal. Normal oil pressure andtemperature indications as per Section 1should be obtained after approximately 5minutes at idle.

2-1-B. HOT WEATHER OPERATIONS

CAUTION

DURING EXTENDED HOVER ATTAKEOFF POWER WITH THE OATABOVE 49.7°C (121.4°F), MONITORTHE ENGINE OIL TEMPERATURE. IFT E M P E R A T U R ER I S E S ABNORMALLY, REDUCE

POWER OR TRANSIT ION TOFORWARD FLIGHT UNTILTEMPERATURE DECREASES.


Each flight should be planned adequately toensure safe operations and to provide pilotwith data to be used during flight.

Check type of mission to be performed anddestination.

Determine that helicopter has adequateperformance to complete mission utilizingappropriate performance charts in Section 4.

Determine that helicopter weight and balancewill be within limits during entire mission.Utilize appropriate weight and balance chartsin Section 5 and limitations in Section 1.

2-3. PREFLIGHT CHECK

Pilot is responsible for determining whetherhelicopter is in condition for a safe flight.Refer to Figure 2-1 for prefl ight checksequence.

NOTE

A preflight check is not intended tobe a detailed mechanical inspection,but simply a guide to help pilot checkcondition of helicopter. It may be ascomprehens ive as condi t ionswarrant at discretion of pilot.

All areas checked shall include av isua l check for ev idence ofcorros ion, par t icu lar ly whenhelicopter is flown near salt water orin areas of high industrial emissions.

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2-4———Rev. 5—19 FEB 2007

2-3-A. BEFORE EXTERIOR CHECK

1. Flight planning — Completed.

2. Publications — Checked.

3. GW and CG — Computed.

4. Helicopter servicing — Completed.

5. Battery — Connected.

2-3-B. EXTERIOR CHECK

2-3-B-1. FUSELAGE — CABIN RIGHT SIDE

WARNING

FAILURE TO REMOVE ROTORT I E D O W N S B E F O R E E N G I N ESTARTING MAY RESULT IN SEVEREDAMAGE AND POSSIBLE INJURY.

1. All main rotor blades — Tiedownsremoved, condition.

2. Right static port — Condition.

3. Cabin doors and hinge bolts —Condition and security.

4. Windows — Condition and security.

5. Landing gear — Condition. Groundhandling wheel removed.

6. Forward and aft crosstube fairings (ifinstalled) — Secured, condition, andaligned.

2-3-B-2. FUSELAGE — CENTER RIGHTSIDE

1. Engine inlet — Condition; removeinlet covers.

2. Cabin roof, transmission cowling,and engine air inlet area — Cleanedof all debris, accumulated snow andice; cowling secured.

3. Forward fairing — Secured.

4. Transmission — Check oil level.Verify actual presence of oil in sightgauge.

5. Transmission oil cooler lines —Condition and security.

6. Transmission mounts — Conditionand security.

7. Main driveshaft — Condition.

8. Access door — Secured.

9. Fuel filler cap — Visually check fuellevel and cap secured.

NOTE

If helicopter is not parked on a levelsurface, fuel sump may not properlydrain contaminants.

10. Fuel sump — Drain fuel sample asfollows:

a. RIGHT and LEFT FUEL BOOST/XFR circuit breaker switches —OFF.

b. BATT switch — BATT (on).

c. FUEL VALVE switch — OFF.

d. FWD and AFT FUEL SUMP drainbuttons — Press, drain sample,then release.

11. Airframe fuel filter — Drain andcheck before first flight of day asfollows:

a. RIGHT and LEFT FUEL BOOST/XFR circuit breaker switches —LEFT and RIGHT (on).

b. FUEL VALVE switch — ON.

c. Fuel filter drain valve — Open,drain sample, then close.

12. Fuel filter test switch — Press andcheck FUEL FILTER caution lightilluminates. Release switch andcheck light extinguishes.


19 FEB 2007—Rev. 5———2-5

13. FUEL VALVE switch — OFF.

14. LEFT and RIGHT FUEL BOOST/XFRcircuit breaker switches — OFF.

15. BATT switch — OFF.

16. Power plant area:

a. Main driveshaft aft flexure —Condition.

b. Engine — Condition, security ofattachments, evidence of oilleakage.

c. Engine mounts — Condition andsecurity.

d. Throttle linkage — Condition,security, and freedom ofoperation.

e. Engine fuel pump — Securityand condition, evidence ofleakage.

f. Hydromechanical unit —Security and condition,evidence of leakage.

g. Hoses and tubing — Chafing,security, and condition.


18. Generator cooling scoop — Clear ofdebris.

19. Oil tank — Leaks, security, and capsecured.

20. Access door — Secured.

21. Aft fairing — Secured.

2-3-B-3. FUSELAGE — AFT RIGHT SIDE

1. Fuselage — Condition.

2. Tail rotor driveshaft cover —Condition and security.

3. Tailboom — Condition.

4. Horizontal stabilizer and positionlight — Condition and security.

2-3-B-4. FUSELAGE — FULL AFT

1. Vertical fin — Condition.

2. Tail rotor guard — Condition andsecurity.

3. Anticollision light — Condition andsecurity of lens.

4. Aft position light — Condition.

5. Tail rotor gearbox — Oil level, leaksand security.

6. Tail rotor — Tiedown removed,condition and free movement.

7. Tail rotor controls — Condition andsecurity.

8. Tail rotor blades:

a. General condition.

b. Tip block — Security and sealintegrity.

c. Internal blade root — Clear ofsnow and ice.

9. Tail rotor yoke — Condition,evidence of static stop contactdamage (deformed static stop yieldindicator).

2-3-B-5. FUSELAGE — AFT LEFT SIDE

1. Tailboom — Condition.

2. Tail rotor driveshaft cover —Condition and security.

3. Horizontal stabilizer area:

a. Horizontal stabilizer — Generalcondition and security ofattachment.

b. Position light — Condition andsecurity.

c. Forward and aft section of leftupper stabilizer support totailboom area — Condition oftailboom.

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2-6———Rev. 5—19 FEB 2007

4. Fuselage — Condition.

5. Forward tail rotor driveshaftcoupling — Condition of splinedadapter.

6. Oil cooler blower shaft hangerbearings — Evidence of greaseleakage and overheating.

7. Oil cooler blower — Clear ofobstructions and condition.

8. Oil cooler — Condition and leaks.

9. Oil cooler blower access door —Secured.

10. Oil tank sight glass — Check oil level.

11. Aft fairing — Secured.

12. Baggage compartment — Cargo tieddown, door secured.

13. Exhaust cover — Removed.

14. Power plant area:

a. Engine — Condition, security ofattachments.

b. Engine mounts — Condition andsecurity.

c. Exhaust stack — Condition andsecurity.

d. Evidence of fuel and oil leaks.

e. Fuel and oil filter bypassindicators — Check retracted.

f. Hoses and tubing for chafingand condition.

g. Pneumatic lines — Conditionand security.

h. Tail rotor driveshaft —Condition of splines andcouplings.

i. Air induction diffuser duct —Condition and security.

j. Rotor brake disc and caliper (ifinstalled) — Condition, securityof attachment and leakage.Ensure brake pads are retractedfrom brake disc.

k. Engine cowling — Secured.

l. Air induction cowling —Secured.

m. Cabin roof, transmissioncowling, engine air inlet area,and plenum — Clear of alldebris, accumulated snow andice; cowling secured.

15. Transmission area:

a. Transmission mounts —Condition and security ofelastomeric mounts.

b. Transmission oil filter — Ensurebypass indicator not extended.

c. Main driveshaft — Condition.

d. Transducers and pressure lines— Condition and security.

e. Access door — Secured.

2-3-B-6. CABIN ROOF

1. Main rotor dampers and fairing —Condition and security.

2. Main rotor hub, yoke and frahm —Condition and security.

3. Main rotor blade and skin —Condition.

4. Pitch horn bearing — Wear andsecurity.

5. Main rotor pitch links — Conditionand security of attachment bolts andlocking hardware.

6. Swashplate assembly — Condition,security of attached controls, andboot condition.

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19 FEB 2007—Rev. 5———2-7

7. Control linkages to swashplate —Condition, security of attachmentbolts and locking hardware.

8. Control tube hydraulics-off balancesprings — Condition and security.

9. Hydraulic reservoir filler cap —Closed and locked.

10. Hydraulic system filters — Bypassindicator retracted.

11. Hydraulic actuators and lines —Condition, security, interference,leakage.

2-3-B-7. FUSELAGE — CABIN LEFT SIDE

1. Forward fairing and access door —Secured.

2. Cabin doors and hinge bolts —Condition and security.

3. Windows — Condition and security.

4. Hydraulic reservoir — Check fluidlevel.

5. Landing gear — Condition andground handling wheel removed.

6. Forward and aft crosstube fairings (ifinstalled) — Secured, condition, andaligned.

7. Left static port — Condition.

2-3-B-8. FUSELAGE — FRONT

1. Exterior surfaces — Condition.

2. Windshield — Condition andcleanliness.

3. Battery and vent lines — Conditionand security.

4. HOUR METER circuit breaker — In.

5. Battery access door — Secured.

6. Pitot tube — Cover removed, clear ofobstructions.

7. External power door — Conditionand security.

8. Landing light lamps — Condition.

9. Antennas — Condition and security.

2-4. INTERIOR AND PRESTARTCHECK

1. Cabin interior — Clean, equipmentsecured.

2. Fire extinguisher — Installed andsecured.

3. Cabin loading — Maintain CG withinlimits.

4. Passenger seat belts — Secured.

5. Copilot seat belt — Secured (if solo).

6. Doors — Secured.

7. Throttle — Closed.

8. LDG LTS switch — OFF.

9. Communications switches — Set.

10. Altimeter — Set.

11. Instruments — Correct indications.

12. Overhead switches — Set:

a. BATT switch — OFF.

b. GEN switch — OFF.

c. PART SEP switch (if installed) —OFF.

d. ANTI COLL LT switch — ANTICOLL LT (on).

e. HYD SYS switch — HYD SYS(on).

f. CABIN LT/PASS switch — OFF.

g. POS LT switch — As desired.

h. DEFOG switch — OFF.

i. PITOT HEATER switch — OFF.


2-8———Rev. 7—30 JUL 2008

j. ENG ANTI ICE switch — OFF.

k. AVIONICS MASTER switch —OFF.

l. HEATER switch (if installed) —OFF.

m. INSTR LT rheostat — OFF.

13. Overhead circuit breaker switches —OFF.

14. Overhead circuit breakers — In.

15. Rotor brake handle (if installed) — Upand latched.

CAUTION

28 VDC GPU SHALL BE 500AMPERES OR LESS TO REDUCERISK OF STARTER DAMAGE FROMOVERHEATING.

16. GPU — Connected (if used).

17. BATT switch — ON for battery start,ON for GPU start, OFF for battery cartstart. Observe the following:

a. Low rotor audio horn activated.

NOTE

With “Ah lers” NR/NP gaugeinstalled, NR/NP needles do not selftest.

b. For 8 seconds,

(1) Trend arcs on LCDinstruments indicate fullscale.

(2) TORQUE and NG digitsdisplay 8188.8.

(3) MGT and FUEL digitsdisplay 81888.

(4) NR and NP needles move to107% and 100%,respectively.

c. After 3 seconds; ENG OUT,FADEC DEGRADE, FADECFAULT, RESTART FAULT, andENGINE OVSPD lightsilluminate with activation ofengine out audio for 3 seconds.

d. ENG OUT light re-illuminateswith reactivation of engine outaudio, after 3 seconds.

18. HORN MUTE button — Press to mute.

19. Caution lights — ENG OUT, XMSNOIL PRESS, RPM, HYDRAULICSYSTEM, GEN FAIL, L/FUEL BOOST,R/FUEL BOOST, L/FUEL XFR, and R/FUEL XFR will be illuminated.

NOTE

L/FUEL XFR and R/FUEL XFR will notbe illuminated when forward fueltank is empty.

20. PEDAL STOP PTT switchannunciator:

Pedals — Centered.

Press — Verify PEDAL STOP cautionand ENGAGED annunciatorilluminated and left pedal travelrestricted.

Release — Verify PEDAL STOPcaution and ENGAGED annunciatorextinguished and both pedals travelunrestricted.

21. Flight controls — Loosen frictions;check travel and verify CYCLICCENTERING light operation; positionfor start. Tighten friction as desired.

22. Throttle — Check freedom of traveland appropriate operation at OFF, I(idle), FLY and MAX positions.Return throttle to OFF position.

NOTE

With INSTR LT rheostat on and CAUTLT switch positioned to DIM, caution


30 JUL 2008—Rev. 7———2-9

lights are dimmed to a fixed intensityand cannot be adjusted by INSTR LTrheostat.

23. INSTR LT rheostat — As desired.

24. CAUT LT switch — As desired.

25. FUEL BOOST/XFR circuit breakerswitches — LEFT (on) and RIGHT(on) and verify all boost and transfercaution lights extinguish.

26. FUEL pressure — Check.

27. CAUTION LT TEST button — Press totest.

28. INSTR CHK button — Press andcheck for exceedances.

29. LCD TEST button — Press to test, ifdesired.

30. FADEC HORN TEST button — Pressto test.

31. FADEC MODE switch — AUTO.

32. FUEL VALVE switch — ON, guardclosed, FUEL VALVE lightilluminates then extinguishes.

33. FUEL QTY — Check TOTAL and FWDtank quantity.

34. OAT/VOLTS display — Check OATand select VOLTS.

CAUTION

ANY ATTEMPT TO START ENGINEWHEN VOLTAGE IS BELOW 24VOLTS MAY RESULT IN A HOTSTART. MONITOR FOR FADECFAILURE. IF FADEC FAILS (FADECFAIL WARNING LIGHT), ABORTSTART BY ROLLING THROTTLE TOCUTOFF AND ENGAGE STARTERTO REDUCE MGT.

2-5. ENGINE START

1. Collective — Full down.

2. Cyclic and pedals — Centered andCYCLIC CENTERING lightextinguished.

NOTE

If throttle is positioned in idle formore than 60 seconds, starterlatching is disabled and throttle mustbe repositioned to cut off and thenback to idle to enable it for another 60seconds.

It is recommended that MGT bebelow 150°C when below 10,000 feetHP or below 65 °C when above 10,000feet HP prior to attempting an enginesta r t . Compl iance w i th th isrecommendation will allow for coolerstarts and reduce potentia l ofreaching hot start abort limits. Referto DRY MOTORING RUN, paragraph2-5-A.

3. Throttle — Idle position.

4. START switch — Momentarily press(hold for approximately 1 second)and observe START and AUTORELIGHT lights are illuminated.

5. MGT — Monitor.

CAUTION

IF MAIN ROTOR IS NOT ROTATINGBY 25% NG, ABORT START BYROLLING THROTTLE TO CUTOFF.ENSURE STARTER HASDISENGAGED WHEN MGTDECREASES BELOW 150°C.

6. START light — Extinguished at 50%NG (starter has disengaged).

7. AUTO RELIGHT light — Extinguishedat 60% NG.

8. ENG and XMSN OIL pressures —Check.

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2-10———Rev. 5—19 FEB 2007

CAUTION

IF ENGINE HAS BEEN SHUT DOWNFOR MORE THAN 15 MINUTES,STABILIZE AT IDLE FOR 1 MINUTEBEFORE INCREASING THROTTLE.

NOTE

During cold temperature operations,normal transmission and engine oilpressure limits may be exceededduring start. Stabilize engine at idleuntil minimum temperature andpressure limits are attained.

9. Idle — 63 ±1% NG.

10. BATT switch — ON (if applicable).

11. GPU — Disconnect and close door (ifapplicable).

12. GEN switch — GEN (on); observeGEN FAIL light extinguishes.

NOTE

Turn generator OFF if ammeterindication drops to zero amps afteran initial full scale indication. Onereset is allowed. RESET generatorand then turn generator back ON(applicable with AMPS/FUEL PSIgauge PN 407-075-024-101 and sub.).Refer to BHT-407-MD-1.

13. Voltmeter — 28.5 ±0.5 volts.

14. FLIGHT INSTR circuit breakerswitches (3) (if installed) — DG, ATTand TURN (on).

NOTE

If dual controls are installed, guardthrottle to prevent inadvertentmanipulation from co-pilot position.

2-5-A. DRY MOTORING RUN

The following procedure is used to reduceresidual MGT to recommended levels forengine start.

1. Throttle — Closed position.

2. START switch — Hold engaged for 15seconds, then release.

Follow ENGINE START procedure, paragraph2-5, once 0% NG is indicated.

2-5-B. ALTERNATE ENGINE START

This procedure may be used in hot and/orhigh altitude environment where aborted hotstarts have been experienced and when priortroubleshooting has not revealed any enginemaintenance issues.



NOTE

It is recommended that MGT bebelow 150°C when below 10,000 feetHP or below 65°C when above 10,000feet HP prior to attempting an enginesta r t . Compl iance w i th th isrecommendation will allow for coolerstarts and reduce potentia l ofreaching hot start abort limits. Referto DRY MOTORING RUN, paragraph2-5-A.

3. Throttle — Closed position.

4. START switch — Hold engaged andobserve START and AUTO RELIGHTlights are illuminated.

a. Throttle — Open to IDLE atapproximately 16% NG.

5. MGT — Monitor.


19 FEB 2007—Rev. 5———2-11

NOTE

Engine will detect l ight off andsmoothly accelerate to idle whilelimiting MGT, if necessary.

a. START switch — Once light offis detected, release.

CAUTION

IF MAIN ROTOR IS NOT ROTATINGBY 25% NG, ABORT START BYROLLING THROTTLE TO CUTOFF.ENSURE STARTER HASDISENGAGED WHEN MGTDECREASES BELOW 150°C.

6. START light — Extinguished at 50%NG (starter has disengaged).

7. AUTO RELIGHT light — Extinguishedat 60% NG.

8. ENG and XMSN OIL pressures —Check.

CAUTION

IF ENGINE HAS BEEN SHUT DOWNFOR MORE THAN 15 MINUTES,STABILIZE AT IDLE FOR 1 MINUTEBEFORE INCREASING THROTTLE.

NOTE

During cold temperature operations,normal transmission and engine oilpressure limits may be exceededduring start. Stabilize engine at idleuntil minimum temperature andpressure limits are attained.

9. Idle — 63 ±1% NG.

10. BATT switch — ON (if applicable).

11. GPU — Disconnect and close door (ifapplicable).

12. GEN switch — GEN (on); observeGEN FAIL light extinguishes.

NOTE

Turn generator OFF if ammeterindication drops to zero amps afteran initial full scale indication. Onereset is allowed. RESET generatorand then turn generator back ON(applicable with AMPS/FUEL PSIgauge PN 407-075-024-101 andsubsequent). Refer to BHT-407-MD-1.

13. Voltmeter — 28.5 ±0.5 volts.

14. FLIGHT INSTR circuit breakerswitches (3) (if installed) — DG, ATTand TURN (on).

NOTE

If dual controls are installed, guardthrott le to prevent inadvertentmanipulation from co-pilot position.

2-6. SYSTEMS CHECK

2-6-A. PRELIMINARY HYDRAULICSYSTEMS CHECK

NOTE

Uncommanded control movement ormotoring with hydraulic system offmay indicate hydraulic systemmalfunction.

1. HYD SYS switch — OFF.

2. HYDRAULIC SYSTEM caution light— Illuminated.

3. HYD SYS switch — HYD SYS (on).

4. HYDRAULIC SYSTEM caution light— Extinguished.


2-12———Rev. 6—20 JUN 2007

2-6-B. FADEC MANUAL CHECK

WARNING

A U T O T O M A N U A L M O D ETRANSITIONS WITH NR/NP AT 100%FLAT PITCH CAN RESULT IN RAPIDN R / N P A C C E L E R A T I O N I NAPPROXIMATELY 7 SECONDS. TOAVOID POSSIBLE OVERSPEEDC O N D I T I O N , P E R F O R M T H EFOLLOWING CHECK AT IDLE (63%NG).

1. Throttle — Idle (63% NG).

2. FADEC MODE switch — MAN.

3. FADEC MANUAL and AUTORELIGHT lights — Illuminated.

4. Check NG stabilized at 75% or less.

5. Throttle — Increase slowly to ensureengine responds, then return to idle.

6. FADEC MODE switch - AUTO.

7. FADEC MANUAL and AUTORELIGHT lights — Extinguished.

2-6-C. ENGINE RUNUP

CAUTION

FAILURE TO SMOOTHLYTRANSITION THE THROTTLE FROMTHE GROUND IDLE POSITION TOTHE FLY POSITION, MAINTAININGTORQUE LESS THAN 40% MAYRESULT IN ENGINE OVERSPEED OROVERTORQUE.

1. Throttle — Increase smoothly to FLYdetent position while maintainingtorque below 40%. Check RPMwarning light extinguished at 95%NR.

2. NR and NP needles — Checkmatching and indicating 100%.

NOTE

Overhead c i rcu i t breakershighlighted with arrow graphic ;are powered through AVIONICSMASTER switch.

3. AVIONICS MASTER switch —AVIONICS MASTER (on).

4. ELT (if installed) — Check forinadvertent transmission.

5. Flight controls — Check freedomwith minimum friction.

6. ENG ANTI ICE switch — ENG ANTIICE (on); check for MGT increase andillumination of ENGINE ANTI-ICElight (if installed).

7. ENG ANTI ICE switch — OFF; checkMGT returns to normal and ENGINEANTI-ICE light (if installed)extinguishes; then ENG ANTI ICE(on) if required.

NOTE

If temperature is below 5°C (40°F)and visible moisture is present, ENGANTI ICE shall be on.

8. PART SEP switch (if installed) — Asrequired.

2-6-D. HYDRAULIC SYSTEMS CHECK

NOTE

Hydraulic systems check is todetermine proper operat ion ofhydraulic actuators for each flightcontrol system. If abnormal forces,unequal forces, control binding, ormotoring are encountered, it may bean indication of a malfunctioningflight control actuator.



20 JUN 2007—Rev. 6———2-13

2. NR — 100% RPM.

3. HYD SYS switch — OFF.

4. HYDRAULIC SYSTEM caution light— Illuminated.

5. Cyclic — Centered.

6. Cyclic control — Check normaloperation by moving cyclic forwardand aft, then left and right(approximately 1 inch). Center cyclic.

7. Collective — Check normaloperation by increasing collectiveslightly (1 to 2 inches). Repeat two tothree times as required. Return to fulldown position.

8. Pedals — Check normal operation bydisplacing pedals slightly (1 inch).

9. HYD SYS switch — HYD SYS (on).

10. HYDRAULIC SYSTEM caution light— Extinguished.

11. Cyclic and collective friction — Setas desired.

2-7. BEFORE TAKEOFF

1. ENG ANTI ICE switch — As required.

2. Light switches — As required.

3. INSTR LT rheostat — As desired.

NOTE

For night flight, it is recommended topoint the map light at the flightinst ruments and set to a lowintensity. Sufficient night lighting willbe provided in the event of aninstrument lighting failure.

4. Radio(s) — Check as required.

5. Flight controls — Position and adjustfrictions for takeoff.

CAUTION

F A I L U R E T O P O S I T I O N A N DMAINTAIN THROTTLE IN FLYDETENT POSIT ION PRIOR TOTAKEOFF AND DURING NORMALFL I GHT OPERATIONS CAN LIMITAVAILABLE ENGINE POWER.

6. Throttle — Open to FLY detentposition. Check 99 to 100% NR/NP.

7. Engine, transmission, and electricalinstruments — Within limits.

8. Flight and navigation instruments —Check.

9. FUEL QTY — Note indication.

10. FUEL QTY FWD TANK button —Press, note fuel remaining in forwardcell.

2-8. TAKEOFF

1. Rear facing seat headrests —Adjusted to proper position.

NOTE

During takeoffs disregard CYCLICCENTERING light and position cyclicas required.

2. Collective — Increase to hover.

3. Directional control — As required tomaintain desired heading.

4. Cyclic — Apply as required toaccelerate smoothly.

5. Increase collective, up to 5% torqueabove hover power, to obtain desiredrate of climb and airspeed. Onceclear of the HV diagram shadedareas, adjust power and airspeed asdesired.


2-14———Rev. 6—20 JUN 2007

6. PEDAL STOP PTT switch — CheckENGAGED annunciator illuminatedabove 55 ±5 KIAS.

2-9. IN-FLIGHT OPERATIONS

1. AIRSPEED — As desired (not toexceed VNE at flight altitude).

CAUTION

A T H I G H P O W E R A N D H I G HA I R S P E E D , C Y C L I C O N L YA C C E L E R A T I O N S A N DMANEUVERING MAYSIGNIFICANTLY INCREASE MGTAND TORQUE WITH NOCOLLECTIVE INPUT. TH ISINCREASE IS MORE RAPID ATLOWER OAT.

NOTE

Pilot shall keep feet on tail rotorpedals at all times. Do not pressPEDAL STOP PTT switch in flight.

2. PEDAL STOP PTT switch — CheckENGAGED annunciator illuminatedabove 55 ±5 KIAS.

3. ENG ANTI ICE and PITOT HEATERswitches — ENG ANTI ICE and PITOTHEATER switches on in visiblemoisture when ambient temperatureis at or below 5°C (40°F).

4. PITOT HEATER — Confirm operation(increase ammeter load).

NOTE

When ENG ANTI ICE switch is in ENGANTI ICE (on), MGT will increase.Monitor MGT when selecting ENGANTI ICE at high power settings.

5. Altimeter — Within limits.

6. FUEL QTY FWD TANK button —Press, note forward fuel tankindication.

NOTE

Full forward fuel tank quanti ty(approximately 256 pounds) will beindicated at approximately 770pounds or greater total fuel. Fuelt ransfer w i l l be complete a tapproximately 195 pounds total fuel.

2-10. DESCENT AND LANDING

NOTE

Large reductions in collective pitchat heavy GW may permit NR toincrease independent of NP (needlessplit). Main rotor may be reengagedwith a smooth increase in collectivepitch.

1. Rear facing seat headrests —Adjusted to proper position.

2. Flight controls — Adjust friction asdesired.

3. Throttle — Fly detent position. Check99 to 100% NP.

4. Flight path — As required for type ofapproach.

5. ENG ANTI ICE — As required.

6. LDG LTS switch — As desired.

NOTE

During run-on or slope landings,disregard CYCLIC CENTERING lightand position cyclic as required. Afterlanding is completed and collectiveis full down, reposition cyclic so thatCYCLIC CENTERING l ight isextinguished.

7. PEDAL STOP PTT switch — CheckENGAGED annunciator extinguishedbelow 50 ±5 KIAS.


20 JUN 2007—Rev. 6———2-15

2-11. ENGINE SHUTDOWN



3. Cyclic friction — Increase so thatcyclic maintains centered position.

4. LDG LTS switch — OFF.

5. Throttle — Reduce to idle stop.Check RPM warning light illuminatedand audio on at 95% NR.

NOTE

If dual controls are installed, guardthrottle to prevent inadvertentmanipulation from co-pilot position.


7. MGT — Stabilize at idle for 2 minutes.

8. ENG ANTI ICE switch — OFF.

9. FLIGHT INSTR circuit breakersswitches (if installed) — OFF.

10. FUEL BOOST/XFR LEFT circuitbreaker switch — OFF.

NOTE

Left fuel boost and transfer pumpswill continue to operate until eitherLEFT FUEL BOOST/XFR circuitbreaker switch (highlighted withyellow border) or FUEL VALVEswitch is positioned to OFF. Thesepumps operate directly from batteryand will not be deactivated whenBATT switch is OFF. Battery powerwill be depleted if both switchesremain on.

11. ELT (if installed) — Check forinadvertent transmission.

12. AVIONICS MASTER switch — OFF.

13. GEN switch — OFF.

14. OVSPD TEST button — If required;press, hold 1 second, and release.

NOTE

Overspeed shutdown test should beaccompl ished on f i rs t eng ineshutdown of the day . ENGINEOVSPD l ight wi l l momentar i lyilluminate in addition to those lightsthat il luminate during a normalshutdown.

15. IDLE REL switch — Press and hold.

CAUTION

POSITIONING THROTTLE OUT OFCUT-OFF DURING NG SPOOL DOWNMAY CAUSE POST ENGINESHUTDOWN FIRE.

16. Throttle — Closed; check MGT andNG decreasing, ENGINE OUTwarning light illuminated and audioon at 55 ±1%.


CAUTION

AVOID RAPID ENGAGEMENT OFROTOR BRAKE IF HELICOPTER ISON ICE OR OTHER SLIPPERY ORLOOSE SURFACE TO PREVENTROTATION OF HELICOPTER.

18. Rotor brake (if installed) — Apply fullrotor brake at or below 40% NR.Return rotor brake handle to stowedposition just prior to main rotorstopping.



2-16———Rev. 7—30 JUL 2008

CAUTION

DO NOT INCREASE COLLECTIVE ORAPPLY LEFT TAIL ROTOR PEDALTO SLOW ROTOR DURINGCOASTDOWN.

20. Pilot — Remain on flight controlsuntil rotor has come to a completestop.

21. All overhead switches, except HYDSYS switch — OFF.

NOTE

Ensure engine ro ta t ion hascomplete ly s topped pr io r topositioning BATT switch to OFF.

22. BATT switch — OFF, with NG at 0%.

CAUTION

APPLICABLE MAINTENANCEACTION MUST BE PERFORMEDPRIOR TO FURTHER FLIGHT IF AFADEC LIGHT HAS ILLUMINATEDDURING THE PREVIOUS FLIGHT ORON ENGINE SHUTDOWN.

NOTE

If shutting down at, or refueling to,between approximately 195 to 213

pounds total fuel quantity, up to 18pounds of fuel may remain in forwardfuel cell as unusable.

2-12. POSTFLIGHT CHECK

If any of following conditions exist:

• Thunderstorms are in local area orforecasted.

• Winds in excess of 35 knots or agust spread of 15 knots exists or isforecasted.

• Helicopter is parked within 150 feetof hovering or taxiing aircraft thatare in excess of basic GW ofhelicopter.

• Helicopter to be left unattended.

Perform following:

1. Install main rotor blade tie-downs.

2. Secure tail rotor loosely to tailboomwith tie-down strap to preventexcessive flapping.

3. Install exhaust cover, engine inletprotective plugs and pitot cover.

NOTE

Refer to BHT-407-MD-1 for additionaltie-down data.


Figure 2-1. Preflight check sequence

17 DEC 2002 2-17/2-18


30 JUL 2008 Rev. 7 3-1/3-2

Section 3EMERGENCY/MALFUNCTION PROCEDURES

3TABLE OF CONTENTS


Introduction ............................................................................................ 3-1 ........... 3-3Definitions .............................................................................................. 3-2 ........... 3-3Engine ..................................................................................................... 3-3 ........... 3-3

Engine Failure .................................................................................... 3-3-A ....... 3-3Engine Restart in Flight..................................................................... 3-3-B ....... 3-4Engine Underspeed ........................................................................... 3-3-C ....... 3-6Engine Overspeed ............................................................................. 3-3-D ....... 3-6Engine Compressor Stall .................................................................. 3-3-E ....... 3-6Engine Hot Start/Shutdown .............................................................. 3-3-F........ 3-7Engine Oil Pressure Low or Fluctuating.......................................... 3-3-G....... 3-7Engine Oil Temperature High ........................................................... 3-3-H ....... 3-7Driveshaft Failure............................................................................... 3-3-J........ 3-8FADEC Failure.................................................................................... 3-3-K ....... 3-8

Fire .......................................................................................................... 3-4 ........... 3-9Engine Fire on Ground ...................................................................... 3-4-A ....... 3-9Engine Fire During Flight .................................................................. 3-4-B ....... 3-9Cabin Smoke or Fumes ..................................................................... 3-4-C ....... 3-9

Tail Rotor ................................................................................................ 3-5 ........... 3-10Complete Loss of Tail Rotor Thrust................................................. 3-5-A ....... 3-10Fixed Pitch Failures ........................................................................... 3-5-B ....... 3-10

Hydraulic System................................................................................... 3-6 ........... 3-11Loss of Hydraulic Pressure .............................................................. 3-6-A ....... 3-11Flight Control Actuator Malfunction ................................................ 3-6-B ....... 3-12

Electrical System ................................................................................... 3-7 ........... 3-12Generator Failure ............................................................................... 3-7-A ....... 3-12Excessive Electrical Load ................................................................. 3-7-B ....... 3-12

Fuel System............................................................................................ 3-8 ........... 3-13Cyclic Jam .............................................................................................. 3-9 ........... 3-13Warning, Caution, and Advisory Lights/Messages ............................ 3-10 ......... 3-14

LIST OF TABLES

Table PageSubject Number Number

Warning (Red) Lights............................................................................. 3-1 ........... 3-15Caution (Amber) and Advisory (White/Green) Lights......................... 3-2 ........... 3-16


30 JUL 2008 Rev. 7 3-3


33-1. INTRODUCTIONFollowing procedures contain indications offailures or malfunctions which affect safety ofcrew, hel icopter, ground personnel orproperty; use of emergency features ofprimary and backup systems; and appropriatewarnings, cautions, and explanatory notes.Tables 3-1 and 3-2 list fault conditions andcorrective actions for warning lights andcaution/advisory lights respectively.

NOTE

All corrective action procedureslisted herein assume pilot gives firstpriority to helicopter control and asafe flight path.

A tripped circuit breaker should notbe reset in flight unless deemednecessary for safe completion of theflight.

If a tripped circuit breaker is deemednecessary for safe completion of theflight, it should only be reset onetime.

Helicopter should not be operated followingany precautionary landing until cause ofmalfunct ion has been determined andcorrective maintenance action taken.

3-2. DEFINITIONSFollowing terms indicate degree of urgency inlanding helicopter.

Fo l lowing terms are used to descr ibeoperating condition of a system, subsystem,assembly, or component.

3-3. ENGINE3-3-A. ENGINE FAILURE

3-3-A-1. ENGINE FAILURE — HOVERING

INDICATIONS:

1. Left yaw.

2. ENGINE OUT and RPM warninglights illuminated.

LAND AS SOON AS POSSIBLE

Land without delay at nearest suitable area (i.e., open field) at which a safe approach and landing is reasonably assured.

LAND AS SOON AS PRACTICAL

Landing site and duration of flight are at discretion of pilot. Extended flight beyond nearest approved landing area is not recommended.

Affected Fails to operate in intended or usual manner.

Normal Operates in intended or usual manner.


3-4 17 DEC 2002

3. Engine instruments indicate powerloss.

4. Engine out audio activated when NGdrops below 55%.

5. NR decreasing with RPM warninglight and audio on when NR dropsbelow 95%.

PROCEDURE:

1. Maintain heading and attitudecontrol.

2. Collective — Adjust to control NRand rate of descent. Increase prior toground contact to cushion landing.

NOTE

Amplitude of collective movement is afunction of height above ground. Anyforward airspeed will aid in ability tocushion landing.

3. Land.

4. Shut down helicopter.

3-3-A-2. ENGINE FAILURE — INFLIGHT

INDICATIONS:

1. Left yaw.

2. ENGINE OUT and RPM warninglights illuminated.

3. Engine instruments indicate powerloss.

4. Engine out audio activated when NGdrops below 55%.

5. NR decreasing with RPM warninglight and audio on when NR dropsbelow 95%.

PROCEDURE:


2. Collective — Adjust as required tomaintain 85 to 107% NR.

NOTE

Mainta in ing NR at h igh end ofoperating range will provide maximumrotor energy to accomplish landing, butwill cause an increased rate of descent.

3. Cyclic — Adjust to obtain desiredautorotative AIRSPEED.

NOTE

Maximum AIRSPEED for steady stateautorotation is 100 KIAS. Minimum rateof descent a irspeed is 55 KIAS.Maximum glide distance airspeed is 80KIAS.

4. Attempt engine restart if amplealtitude remains. (Refer to ENGINERESTART, paragraph 3-3-B).


6. At low altitude:

a. Throttle — Closed.

b. Flare to lose airspeed.

7. Apply collective as flare effectdecreases to further reduce forwardspeed and cushion landing. Uponground contact, collective shall bereduced smoothly while maintainingcyc l ic in neut ra l or cente redposition.

8. Complete helicopter shutdown.

3-3-B. ENGINE RESTART IN FLIGHT

An engine restart may be attempted in flight iftime and altitude permit.


17 DEC 2002 3-5

CAUTION

IF CAUSE OF FAILURE IS OBVIOUSLYMECHANICAL, AS EVIDENCED BYABNORMAL METALLIC OR GRINDINGSOUNDS, DO NOT ATTEMPT ARESTART.

3-3-B-1. RESTART − AUTOMATIC MODE

l PROCEDURE (NO RESTART FAULT ORFADEC MANUAL LIGHTS ILLUMINATED):

1. Collective — Adjust to maintain 85 to107% NR.

2. AIRSPEED — Adjust as desired.

NOTE

Minimum rate of descent airspeed of 55KIAS and minimum NR will allow pilotmore time for air restart.

3. FUEL VALVE switch — ON.

4. Throttle — Cutoff.

5. START switch — Hold to startposition (start will latch after throttleis placed to idle).

6. NG — Between 12% and 50%.

7. Throttle — Idle.

8. MGT — Monitor.

9. Throttle — Advance smoothly to FLYdetent position.

If restart is unsuccessful, abort start andsecure engine as follows:



12. Accomplish autorotative descentand landing.

3-3-B-2. RESTART — MANUAL MODE

RESTART FAULT OR FADEC MANUALLIGHTS ILLUMINATED.

l PROCEDURE:

1. Collective — Adjust to maintain 85 to107% NR.

2. AIRSPEED — Adjust as desired.

NOTE

Minimum rate of descent airspeed of 55KIAS and minimum NR will allow pilotmore time for air restart.



5. FUEL VALVE switch — ON.

6. START switch — Hold to startposition (starter will not latch).

7. NG — 12%.

8. Throttle — Slowly advance out ofcutoff and stop advancing throttle atlight off.

9. MGT — Allow to peak.

10. Throttle — Increase fuel flow bymodulating throttle to maintain MGTwithin limits.

11. START switch — Release at 50% NG.

12. Throttle — Advance smoothly andmodulate to 100% NP.

If restart is unsuccessful, abort start andsecure engine as follows:



3-6 17 DEC 2002


15. Accomplish autorotative descentand landing.

3-3-C. ENGINE UNDERSPEED

NO CAUTION/WARNING/ADVISORY LIGHTSILLUMINATED.

INDICATIONS:

1. Decrease in NG.

2. Subsequent decrease in NP.

3. Possible decrease in NR.

4. Decrease in TRQ.

PROCEDURE:


2. Throttle — Confirm in FLY detentposition.

3. Throttle — Position throttle to theapproximate bezel position thatcoincides with the gauge indicatedNG.


5. NR — Maintain 95 to 100% withthrottle and collective.

6. Land as soon as practical.

3-3-D. ENGINE OVERSPEED

(NO CAUTION/WARNING/ADVISORY LIGHTSILLUMINATED)

INDICATIONS:

1. Increase in NR.

2. Increase in NP.

3. Increase in NG.

4. Increase in TRQ.

PROCEDURE:

1. Throttle — Retard.

2. NG or NP — Attempt to stabilize withthrottle and collective.


4. NR — Maintain 95 to 100% withthrottle and collective.

CAUTION

IF UNABLE TO MAINTAIN NR, NP, NG,OR MGT, PREPARE FOR A POWEROFF LANDING BY LOWERINGCOLLECTIVE AND SHUTTING DOWNENGINE.

3-3-E. ENGINE COMPRESSOR STALL

INDICATIONS:

1. Engine pops.

2. High or erratic MGT.

3. Decreasing or erratic NG or NP.

4. TRQ oscillations.

PROCEDURE:

1. Collective — Reduce power,maintain slow cruise flight.

2. MGT and NG — Check for normalindications.

3. ENG ANTI ICE switch — ON.

4. PART SEP switch (if installed) — ON.


17 DEC 2002 3-7

5. HEATER switch (if installed) — ON.

NOTE

Severity of compressor stalls willdictate if engine should be shut downand treated as an engine failure. Violentstalls can cause damage to engine anddrive system components, and must behandled as an emergency condition.Stalls of a less severe nature (one ortwo low intensity pops) may permitcontinued operation of engine at areduced power leve l , avo id ingcondition that resulted in compressorstall.

If pilot elects to continue flight:

6. Collective — Increase slowly toachieve desired power level.

7. MGT and NG — Monitor for normalresponse.


If pilot elects to shut down engine:

9. Enter autorotation.




13. Cyclic — Adjust as required tomaintain desired AIRSPEED.

14. Prepare for power-off landing.

3-3-F. ENGINE HOT START/SHUTDOWN

INDICATIONS:

1. Excessive MGT.

2. Visible smoke or fire.

PROCEDURE:



NOTE

Starter will remain engaged until MGTdecreases to 150 °C and thenautomatically disengage. Starter maybe manually engaged by holdingSTARTER switch forward.

3. STARTER switch — Ensure starter ismotoring engine until MGT stabilizesat normal temperature.

4. Shut down helicopter.

3-3-G. ENGINE OIL PRESSURE LOW OR FLUCTUATING

INDICATIONS:

1. Engine oil pressure below minimum.

2. Engine oil pressure fluctuatingabnormally.

PROCEDURE:

1. Engine oil pressure and temperature— Monitor.


3-3-H. ENGINE OIL TEMPERATURE HIGH

INDICATIONS:

1. Engine oil temperature increasingabove normal.

2. Engine oil temperature abovemaximum.

PROCEDURE:

Land as soon as practical.


3-8 17 DEC 2002

3-3-J. DRIVESHAFT FAILURE

WARNING

FAILURE OF MAIN DRIVESHAFT TOTRANSMISSION WILL RESULT INCOMPLETE LOSS OF POWER TO MAINROTOR. ALTHOUGH COCKPITINDICATIONS FOR A DRIVESHAFTFAILURE ARE SIMILAR TO AN ENGINEOVERSPEED, IT IS IMPERATIVE THATA U T O R O T A T I V E F L I G H TPROCEDURES BE ESTABLISHEDIMMEDIATELY. FAILURE TO REACTIMMEDIATELY TO LOW RPM AUDIO,RPM LIGHT AND NP/NR TACHOMETERCAN RESULT IN LOSS OF CONTROL.

INDICATIONS:

1. Left yaw

2. Rapid decrease in NR

3. Rapid increase in NP

4. LOW RPM audio horn

5. Illumination of RPM light

6. Possible increase in noise level dueto overspeed ing engine anddriveshaft breakage.

NOTE

Engine overspeed trip system willactivate at 118.5% NP causing fuel flowto go to minimum. After initialoverspeed, FADEC will adjust fuel flowto maintain engine at 100% NP.

PROCEDURE:



NOTE

Minimum rate of descent airspeed is 55KIAS. Maximum glide distanceairspeed is 80 KIAS.

3. Cyclic — Adjust to obtain desiredautorotative airspeed.

NOTE

To maintain tail rotor effectiveness donot shutdown engine.

4. Landing — Complete autorotativelanding.

5. Complete helicopter shutdown.

3-3-K. FADEC FAILURE

NOTE

Takeoff power may not be available inthe MAN mode. Maximum continuouspower will be available for all ambientconditions.

INDICATIONS

1. FADEC fail audio activated.

2. FA D E C FA I L w a r n i n g l i g h tilluminated.

3. FADEC MANUAL caution l ightilluminated.

4. AUTO RELIGHT advisory l ightilluminated.

5. FADEC MODE switch MAN lightilluminated.


30 JUL 2008 Rev. 7 3-9

PROCEDURE:

WARNING

WITHIN 2 TO 7 SECONDS AFTERTHE FADEC FAIL WARNING NR/NPMAY INCREASE RAPIDLY,REQUIRING POSITIVE MOVEMENTSOF COLLECTIVE AND THROTTLETO CONTROL NR.

1. Throttle — If time permits, matchthro t t le beze l pos i t ion to NGindication.

2. NR/NP — Maintain 95 to 100% withcollective and throttle.

3. FADEC MODE switch — Depress onetime, muting FADEC fail audio.

NOTE

Depressing FADEC MODE switchone time, will only mute FADEC failaudio. This step should not beaccomplished until pilot is firmlyestablished in MAN control.


5. Normal shutdown if possible.

3-4. FIRE3-4-A. ENGINE FIRE ON GROUND

INDICATIONS:

1. Smoke

2. Fumes

3. Fire

PROCEDURE:





5. Rotor brake (if installed) — Engage.

6. Exit helicopter.

3-4-B. ENGINE FIRE DURING FLIGHT

INDICATIONS:

1. Smoke

2. Fumes

3. Fire

PROCEDURE:

1. Inflight — Immediately enterautorotation.



4. If time permits, FUEL BOOST/XFRcircuit breaker switches — OFF.

5. Execute autorotative descent andlanding.


NOTE

Do not restart engine until correctivemaintenance has been performed.

3-4-C. CABIN SMOKE OR FUMES

INDICATIONS:

1. Smoke

2. Fumes


3-10 17 DEC 2002

PROCEDURE:

1. Inflight — Start descent

2. AIR COND BLO switch (if installed)— OFF

3. HEATER switch (if installed) — OFF

4. All vents — Open

5. Side windows — Open

If time and altitude permits:

6. Source — Attempt to identify andsecure.

7. If source is identified and smokeand/or fumes still persist — Land assoon as possible.

8. If source is identified and smokeand/or fumes are cleared — Land assoon as practical.

3-5. TAIL ROTOR

There is no single emergency procedure forall types of antitorque malfunctions. One keyto a pilot successfully handling a tail rotoremergency lies in the abil ity to quicklyrecognize the type of malfunction that hasoccurred.

3-5-A. COMPLETE LOSS OF TAIL ROTOR THRUST

This is a situation involving a break in drivesystem (e.g., severed driveshaft), wherein tailrotor stops turning and delivers no thrust.

INDICATIONS:

1. Uncontrollable yawing to right (leftside slip).

2. Nose down tucking.

3. Possible roll of fuselage.

NOTE

Severity of initial reaction of helicopterwill be affected by AIRSPEED, CG,power being used, and HD.

PROCEDURE:

3-5-A-1. HOVERING

Close thrott le and perform a hoveringautorotation landing. A slight rotation can beexpected on touchdown.

3-5-A-2. IN-FLIGHT

Reduce throttle to idle, immediately enterautorotat ion, and maintain a minimumAIRSPEED of 55 KIAS during descent.

NOTE

When a suitable landing site is notavailable, vertical f in may permitcontrolled flight at low power levelsand sufficient AIRSPEED. During finalstages of approach, a mild flare shouldbe executed, making sure all power torotor is off. Maintain helicopter in aslight flare and smoothly use collectiveto execute a soft, slightly nose-highlanding. Landing on aft portion of skidswill tend to correct side drift. Thistechnique will, in most cases, result ina run-on type landing.

CAUTION

IN A RUN-ON TYPE LANDING AFTERTOUCHING DOWN, DO NOT USECYCLIC TO REDUCE FORWARDSPEED.

3-5-B. FIXED PITCH FAILURES

This is a situation involving inability tochange tail rotor thrust (blade angle) with anti-torque pedals.


17 DEC 2002 3-11

INDICATIONS:

1. Lack of directional response.

2. Locked pedals.

NOTE

If pedals cannot be moved with amoderate amount of force, do notattempt to apply a maximum effort,since a more serious malfunction couldresult. If helicopter is in a trimmedcondition when malfunction occurs,TRQ and AIRSPEED should be notedand helicopter flown to a suitablelanding area. Certain combinations ofTRQ, NR, and AIRSPEED will correct ayaw attitude, and these combinationsshould be used to land helicopter.

PROCEDURE:

NOTE

Pull pedal stop emergency release toensure pedal stop is retracted.

3-5-B-1. HOVERING

Do not close throttle unless a severe rightyaw occurs. If pedals lock in any position at ahover, land ing f rom a hover can beaccomplished with greater safety underpower-controlled flight rather than by closingthrottle and entering autorotation.

3-5-B-2. IN-FLIGHT — LEFT PEDAL APPLIED

In a high power condition, helicopter will yawto left when power is reduced. Power andAIRSPEED should be adjusted to a valuewhere a comfortable yaw angle can bemaintained. If AIRSPEED is increased, verticalf in wi l l become more ef fect ive and anincreased left yaw attitude will develop. Toaccomplish landing, establish a power-on

approach with sufficiently low AIRSPEED(zero if necessary) to attain a rate of descentwith a comfortable sideslip angle. (A decreasein NP decreases ta i l ro tor thrust . ) Ascollective is increased just before touchdown,left yaw will be reduced.

3-5-B-3. IN-FLIGHT — RIGHT PEDAL APPLIED

In cruise flight or reduced power situation,helicopter will yaw to right when power isincreased. A low power, run-on type landingwill be necessary by gradually reducingthrottle to maintain heading while addingcollective to cushion landing. If right yawbecomes excessive, close throttle completely.

3-6. HYDRAULIC SYSTEM

3-6-A. LOSS OF HYDRAULIC PRESSURE

INDICATIONS:

1. HYDRAULIC SYSTEM caution lightilluminated.

2. Grinding or howling noise frompump.

3. Increase in force required to moveflight controls.

4. Feedback forces may be evidentduring flight control movement.

PROCEDURE:

1. Reduce AIRSPEED to 70 to 100KIAS.

2. HYD SYSTEM circuit breaker — Out.If hydraulic power is not restored,push breaker in.

3. HYD SYS switch — HYD SYS; OFF ifhydraulic power is not restored.


3-12 17 DEC 2002

4. For extended flight set comfortableAIRSPEED, up to 120 KIAS, tominimize control forces.


6. A run-on landing at effectivet rans la t iona l l i ft speed(approx imate ly 15 knots ) isrecommended.

3-6-B. FLIGHT CONTROL ACTUATOR MALFUNCTION

An actuator hardover can occur in any flightcontrol axis, but a cyclic cam jam will onlyoccur in the fore and aft axis. An actuatorhardover is manifested by uncommandedmovements of one or two flight controls. Iftwo controls move, the pilot will find one ofthese controls will require a higher thannorma l contro l fo rce to oppose themovement. This force cannot be “trimmed” tozero without turning the HYD SYS switch OFF.Once the hydraulic boost is OFF, the forceson the affected flight control will be similar tothe “normal” hydraulic off forces.

INDICATIONS:

1. Uncommanded flight controlmovements

2. High flight control forces to opposemovement in one axis

3. Feedback forces only in affectedflight control axis

4. Flight control forces normal inunaffected axis

PROCEDURE:

1. Attitude — Maintain

2. HYD SYS switch — OFF

3. AIRSPEED — Set to 70 to 100 KIAS

4. Land as soon as possible usingprocedure from paragraph 3-6-A

3-7. ELECTRICAL SYSTEM

3-7-A. GENERATOR FAILURE

INDICATIONS:

1. GEN FAIL caution light illuminated.

2. AMPS indicates 0.

3. Voltmeter — Approximately 24 volts

PROCEDURE:

1. GENERATOR FIELD andGENERATOR RESET c i rcu i tbreakers — Check in.

2. GEN switch — RESET; then GEN.

3. If power is not restored, place GENswitch to OFF; land as soon aspractical.

NOTE

With generator OFF, a fully chargedbattery will provide approximately 21minutes of power for basic helicopterand one VHF COMM radio (35 minuteswith optional 28 ampere/hour battery).

3-7-B. EXCESSIVE ELECTRICAL LOAD

INDICATIONS:

1. AMPS indicates excessive load.

2. Smoke or fumes.

PROCEDURE:




19 FEB 2007 Rev. 5 3-13

3. FUEL BOOST/XFR LEFT circuitbreaker switch — LEFT (on).

WARNING

PRIOR TO BATTERY DEPLETION,ALTITUDE MUST BE REDUCEDBELOW 8000 FEET HP (JET A) OR 4000FEET HP (JET B). UNUSABLE FUELMAY BE AS HIGH AS 150 POUNDSAFTER THE BATTERY IS DEPLETEDDUE TO INABILITY TO TRANSFERFUEL FROM FORWARD CELLS.

NOTE

With battery and generator OFF, an80% charged battery will operate leftfuel boost pump and left fuel transferpump for approximately 1.7 hours (2.8hours with optional 28 ampere/hourbattery).

4. Airspeed — 60 KIAS or less.

NOTE

Pedal stop disengages with loss ofelectrical power.


NOTE

When throttle is repositioned to the idlestop (during engine shutdown) thePMA will go offline and the engine mayflame out.

3-8. FUEL SYSTEMDUAL FUEL TRANSFER FAILURE

INDICATIONS:

1. L/FUEL XFR and R/FUEL XFRcaution lights illuminate.

2. Last 150 pounds of fuel in forwardcell may not be usable.

3. Fuel will stop transferring fromforward to a ft fue l ce l l a tapproximately 345 pounds totalindicated fuel.

PROCEDURE:

1. LEFT and RIGHT FUEL BOOST/XFRcircuit breaker switches — CheckON.

2. Determine FUEL QTY in forward cell.

3. Subtract quantity of fuel trapped inforward cell from total to determineusable fuel remaining.

4. Plan landing accordingly.

3-9. CYCLIC CAM JAMA cyclic cam jam can only occur in the foreand aft axis, whereas, an actuator hardovercan occur in any flight control axis. A cycliccam jam is manifested when a commandedcontrol movement requires a higher thannormal fore and aft spring force. The force feltwhen moving the cyclic fore and aft with acam jam is the result of overriding a springcapsule.

INDICATIONS:

1. High (approximately 15 pounds) foreand aft cyclic control forces.

2. Normal pedal, collective and lateralcyclic control forces.

PROCEDURE:

1. Helicopter pitch attitude — Maintainnormal pitch attitudes with forwardor aft cyclic force.


3-14 17 DEC 2002

CAUTION

DO NOT TURN HYDRAULIC BOOSTOFF


3-10. WARNING, CAUTION, AND ADVISORY LIGHTS/MESSAGES

Red warn ing l ights /messages , fau l tcondit ions, and corrective actions arepresented in Table 3-1.

Amber caution and White advisory lights/messages and correc t ive ac t ions arepresented in Table 3-2.


17 DEC 2002 3-15

Table 3-1. Warning (red) lights

PANELWORDING FAULT CONDITION CORRECTIVE ACTION

BATTERY HOT Battery overheating. Turn BATT switch OFF and land as soon as practical. If BATTERY RLY light illuminates, turn GEN switch OFF if conditions permit. Land as soon as possible.

ENGINE OUT NG less than 55 ± 1% and/or FADEC senses ENGINE OUT.

Verify engine condition. Accomplish engine failure procedure.

ENGINE OVSPD NG greater than 110% or NP versus TORQUE is above maximum continuous limit (102.4% NP at 100% TORQUE to 108.6% NP at 0% TORQUE).

Adjust throttle and collective as necessary. Determine if engine is controllable, if not shut down. Maintenance action required before next flight.

FADEC FAIL (During start)

FADEC has detected a serious malfunction.

Close throttle immediately. Engage starter to reduce MGT. Applicable maintenance action required prior to next flight.

FADEC FAIL (Inflight) FADEC has detected a malfunction and an overspeed may occur 2 to 7 seconds following activation of FADEC fail horn and illumination of FADEC FAIL warning light. Engine may underspeed significantly prior to overspeed. Any other FADEC related lights may be illuminated.

Accomplish FADEC FAILURE procedure, paragraph 3-3-K. Applicable maintenance action required prior to next flight.

RPM (with low RPM audio)

NR below 95%. Reduce collective and ensure throttle is in FLY detent position. Light will extinguish and audio will cease when NR increases above 95%.

RPM (without audio) NR above 107%. Increase collective and/or reduce severity of maneuver. Light will extinguish when NR decreases below 107%.

XMSN OIL PRESS Transmission oil pressure is below minimum.

Reduce power; verify fault with gage. Land as soon as possible.

XMSN OIL TEMP Transmission oil temperature is at or above red line.

Reduce power; verify fault with gage. Land as soon as practical.


3-16 17 DEC 2002

Table 3-2. Caution (amber) and advisory (white/green) lights


AUTO RELIGHT (white) Engine igniter is operating. None.

NOTE

AUTO RELIGHT light will beilluminated when ignition system isactivated.Ignition system isactivated:

1 - during start sequence

2 - in MANUAL mode with NGabove 55%

3 - with FADEC detection of engineout condition with NG above 50%.

BAGGAGE DOOR Baggage compartment door not securely latched.

Close door securely before flight. If light illuminates during flight, land as soon as practical.

BATTERY RLY Battery relay has malfunctioned to closed (ON) position with BATT switch OFF. Battery is still connected to DC BUSS.

If BATTERY HOT light is illuminated, turn GEN switch OFF if conditions permit. Land as soon as possible.

CHECK INSTR TRQ, MGT, or NG is about to or has detected an exceedance. Flashing LCD trend arc and digital display indicates impending exceedance. Letter E in digital display indicates an exceedance has occurred.

Reduce engine power if possible. Press INSTR CHK button to display magnitude of exceedance. Refer to BHT-407-MD-1.

CYCLIC CENTERING Cyclic stick is not centered. Reposition cyclic stick to center position to extinguish CYCLIC CENTERING light.

ENGAGED Information system status. None.

ENGINE ANTI-ICE (white)

ANTI-ICE switch ON. Engine receiving anti-icing air.

If light (if installed) remains illuminated with ENGINE ANTI-ICE switch OFF, avoid operations requiring maximum power.

ENGINE CHIP Ferrous particles in engine oil. Land as soon as possible.


17 DEC 2002 3-17

FADEC DEGRADED (Inflight)

FADEC ECU operation is degraded which may result in NR droop, NR lag, or reduced maximum power capability.

Remain in AUTO mode. Fly helicopter smoothly and nonaggressively. Land as soon as practical.

NOTE

It may be necessary to use FUELVALVE switch to shut down engineafter landing.

Applicable maintenance action required prior to next flight.

FADEC DEGRADED (With engine shutdown)

FADEC ECU has recorded a fault during previous flight or a current fault has been detected.

Position throttle to idle; if light extinguishes, fault is from previous flight. Applicable maintenance action required prior to next flight.

FADEC FAULT PMA and or MGT, NP or NG automatic limiting circuit(s) not functional.

Remain in AUTO mode. Land as soon as practical. Applicable maintenance action required prior to next flight.

FADEC MANUAL FADEC is operating in MANUAL mode. No automatic governing is available. AUTO RELIGHT light will be illuminated.

Fly helicopter smoothly and nonaggressively. Maintain NR with coordinated throttle and collective movements. Land as soon as practical.

FLOAT ARM FLOAT ARM switch is ON. Float inflation solenoid is armed.

Normal operation for takeoff and landing over water. FLOAT ARM switch — OFF. If light remains illuminated, FLOATS circuit breaker — Out. Land as soon as practical.

NOTE

With float inflation solenoid armed,flight should not exceed 60 KIAS and500 feet AGL.

FLOAT TEST (green) Float system in test mode. None.

Table 3-2. Caution (amber) and advisory (white/green) lights (Cont)



3-18 Rev. 5 19 FEB 2007

FUEL FILTER Airframe fuel filter in impending bypass.

Land as soon as practical. Clean before next flight.

FUEL LOW 100 ±10 pounds of fuel remain in aft tank.

Verify FUEL QTY. Land as soon as practical.

R/FUEL BOOST Right fuel boost pump has failed. If practical, descend below 8000 feet HP if fuel is Jet A or 4000 feet HP if fuel is Jet B to prevent fuel starvation if other fuel boost pump fails or has low output pressure. Land as soon as practical.

WARNING

IF BOTH FUEL BOOST PUMPS FAIL,ALTITUDE MUST BE REDUCED TOBELOW 8000 FEET HP (JET A) OR4000 FEET Hp (JET B). LAND ASSOON AS POSSIBLE.

L/FUEL BOOST Left fuel boost pump has failed. If practical, descend below 8000 feet HP if fuel is Jet A or 4000 feet HP if fuel is Jet B to prevent fuel starvation if other fuel boost pump fails or has low output pressure. Land as soon as practical.

FUEL VALVE Fuel valve position differs from FUEL VALVE switch indication or FUEL VALVE circuit breaker out.

Check FUEL VALVE circuit breaker in. Land a soon as practical. If on ground, cycle FUEL VALVE switch.

L/FUEL XFR Left fuel transfer pump has failed.





19 FEB 2007 Rev. 5 3-19

CAUTION

IF BOTH FUEL TRANSFER PUMPSFAIL, UNUSABLE FUEL MAY BE ASHIGH AS 150 POUNDS DUE TOINABILITY TO TRANSFER FUELFROM FORWARD CELL. LAND ASSOON AS PRACTICAL.

NOTE

Under normal fuel transferconditions, helicopters S/N 53000through 53174 L/FUEL XFR andR/FUEL XFR lights will illuminatefor 2.5 minutes and thenextinguish. This indicatestransfer is complete and transferpumps have been automaticallyturned off. Helicopters S/N 53175and subsequent inhibitillumination of the lights.

R/FUEL XFR Right fuel transfer pump has failed.


GEN FAIL Generator not connected to DC BUSS.

Verify fault with AMPS gauge. GEN switch — RESET, then ON. If GEN FAIL light remains illuminated, GEN switch — OFF. Land as soon as practical.

HEATER OVERTEMP An overtemp condition has been detected by a temperature probe either under pilot seat, copilot seat, or in vertical tunnel.

Turn HEATER switch OFF immediately.




3-20 17 DEC 2002

HYDRAULIC SYSTEM Hydraulic pressure below limit. Verify HYD SYS switch position. Accomplish hydraulic system failure procedure (refer to paragraph 3-6).

LITTER DOOR Litter door not securely latched. Close door securely before flight. If light illuminates during flight, land as soon as practical.

PEDAL STOP Pedal Restrictor Control Unit has detected a failure of part of system.

VNE — 60 KIAS.

PEDAL STOP emergency release — Pull.


RESTART FAULT(white)

FADEC ECU has detected a fault which will not allow engine to be restarted in AUTO mode.

Remain in AUTO mode. Plan landing site accordingly.

Applicable maintenance action required prior to next flight.

NOTE

When throttle is repositioned to idlestop (during engine shutdown) thePMA will go offline and engine mayflameout.

START(white)

Start relay is in START mode. If START switch has not been engaged and there is zero indication on AMPS gage; START relay has malfunctioned and helicopter is on battery power. START circuit breaker — Out. Land as soon as practical.

T/R CHIP Ferrous particles in tail rotor gearbox oil.

Land as soon as possible.

XMSN CHIP Ferrous particles in transmission oil.

Land as soon as possible.




17 DEC 2002 4-1/4-2

Section 4PERFORMANCE

4

Paragraph NumberPage

Subject

TABLE OF CONTENTS

INTRODUCTION ......................................................................... 4-1.................... 4-3POWER ASSURANCE CHECK.................................................. 4-2.................... 4-3DENSITY ALTITUDE................................................................... 4-3.................... 4-4HEIGHT – VELOCITY ENVELOPE............................................. 4-4.................... 4-4HOVER CEILING......................................................................... 4-5.................... 4-4NOT USED .................................................................................. 4-6.................... 4-5CLIMB AND DESCENT............................................................... 4-7.................... 4-5

CLIMB................................................................................... 4-7-A................ 4-5 AUTOROTATION ................................................................. 4-7-B................ 4-6

AIRSPEED CALIBRATION......................................................... 4-8.................... 4-6NOT USED .................................................................................. 4-9.................... 4-6NOISE LEVELS........................................................................... 4-10.................. 4-6

FAR PART 36 STAGE 2 NOISE LEVEL................................. 4-10-A.............. 4-6CANADIAN AIRWORTHINESS MANUAL CHAPTER 516 AND ICAO ANNEX 16 NOISE LEVEL............................. 4-10-B.............. 4-6

Figure Title Number Number

Page

LIST OF FIGURES

Power assurance check ............................................................ 4-1.................... 4-7Density altitude .......................................................................... 4-2.................... 4-8Altitude vs gross weight for height – velocity diagram ......... 4-3.................... 4-9Height – velocity diagram ......................................................... 4-4................... 4-10Hover ceiling wind accountability chart – below 14,000 feet HD................................................................ 4-5................... 4-11Hover ceiling wind accountability chart – between 14,000 and 17,000 feet HD ........................................ 4-5A ................ 4-12Hover ceiling wind accountability chart – above 17,000 feet HD ............................................................... 4-5B ................ 4-13Hover ceiling IGE ....................................................................... 4-6................... 4-14Hover ceiling OGE ..................................................................... 4-7................... 4-22Rate of climb – takeoff power ................................................... 4-8................... 4-30Rate of climb – maximum continuous power.......................... 4-9................... 4-40Autorotation glide distance ...................................................... 4-10................. 4-50Airspeed installation correction ............................................... 4-11................. 4-51


17 DEC 2002 4-3


44-1. INTRODUCTION

Performance data presented herein areder ived f rom eng ine manufacture 'sspec i f ica t ion power fo r eng ine lessinstallation losses. These data are applicableto basic helicopter without any optionalequipment that would appreciably affect lift,drag, or power available.

4-2. POWER ASSURANCE CHECK

A Power assurance check chart (Figure 4-1) isprovided for Rolls-Royce model 250-C47Bengine. This chart indicates maximumal lowable MGT for an engine meet ingminimum Rolls-Royce specification. Enginemust develop required torque withoutexceeding chart MGT in order to meetperformance data contained in this manual.

Figure 4-1 may be used to periodicallymonitor engine performance.

To perform power assurance check, turn offall sources of bleed air, including ENGINEANTI-ICING. Establish level f l ight at anAIRSPEED of 85 to 105 K IAS or V N E ,whichever is lower. Check may also beconducted in a hover pr ior to takeoff ,depending on ambient conditions and grossweight.

Record following information from cockpitinstruments:

EXAMPLE:

SOLUTION:

Enter Power assurance check chart atobserved TORQUE (70%), proceed verticallydown to intersect HP (6000 feet), followhorizontally to intersect indicated OAT (10°C),then drop vert ical ly to read maximumallowable MGT.

If actual MGT is less than or equal to chartMGT, engine performance equals or exceedsminimum specification and performance datacontained in this manual can be achieved.

If actual MGT is greater than chart MGT,engine performance is less than minimumspecificat ion and al l performance datacontained in this manual cannot be achieved.Refer to appropriate maintenance manual todetermine cause of low power (high MGT).

NOTE

Chart may also be used to determineminimum specification power foractual MGT. Using above example,enter chart at actual MGT (675°C,proceed up to OAT (10°C), across to HP(6000 feet), and up to read minimumtorque available (70%). If actual poweris equal to or greater than chart torque,engine performance equals or exceedsmin imum spec i f ica t ion andperformance data contained in thismanual can be achieved. If actualtorque indication is less than charttorque, engine performance is lessthan minimum specification and allperformance in this manual cannot beachieved. Re fer to appropr ia temaintenance manual to determinecause of low power.

HP 6000 feet

OAT 10°C

MGT Actual reading

TORQUE 70%


4-4 17 DEC 2002

4-3. DENSITY ALTITUDE

A Dens i ty a l t i tude and temperatureconversion chart (Figure 4-2) is provided toaid in calculat ion of performance andlimitations. HD is an expression of density ofair in terms of height above sea level; hence,the less dense the air, the higher the HD. Forstandard conditions of temperature andpressure, HD is same as HP. As temperatureincreases above standard for an altitude, HDwill also increase to values higher than HP.Figure 4-2 expresses HD as a function of HPand temperature.

Density altitude chart also includes theinverse of the square root of the density ratio

(1 / ) , which is used to calculate trueairspeed by the following relation:

KTAS = KCAS×1/

EXAMPLE:

If ambient temperature is -15°C and HP is 7000

feet, find HD, 1/ , and true airspeed for 100KCAS.

SOLUTION:

Enter bottom of chart at -15°C.

Move vertically upward to 7000 feet HP line.

From this point, move horizontally toleft and read HD of 5000 feet, move

horizontally to right and read 1/ =1.08.

True airspeed = KCAS × 1/ = 100 ×1.08 = 108 KTAS.

4-4. HEIGHT – VELOCITYENVELOPE

The Height – Velocity envelope charts(Figures 4-3 and 4-4) define conditions fromwhich a safe landing can be made on asmooth, level, firm surface; following anengine failure. The Height – Velocity diagram

(Figure 4-4) is valid only when helicoptergross weight does not exceed limits of theAltitude vs Gross Weight for Height – Velocitydiagram (Figure 4-3). Four envelopes (GrossWeight Regions) are specified. Each GrossWeight Region applies for all gross weightswithin its boundaries. No interpolation isallowed.

For a given ambient outside air temperature,pressure altitude, and gross weight, theappropriate limiting envelope (Region A, B, C,or D) can be determined. Using Figure 4-3(Altitude VS Gross Weight), move upwardvertically from entry OAT to pressure altitude.From that point, move right horizontally tode termine the correct weigh t reg ion .(Examples: 15°C at Sea Level at 5000 poundsGW = Region B, and 30°C at 2000 feetpressure altitude at 5000 pounds GW =Region D) Once the correct weight region hasbeen dete rmined (A , B , C , or D ) , thecorresponding Avoid area is selected fromFigure 4-4 (Height – Velocity diagram).

4-5. HOVER CEILING

NOTE

Hover performance charts are basedon 100% ROTOR RPM.

Satisfactory stability and control have beendemonstrated in each area of the Hoverceiling charts with winds as depicted on theHover ceiling wind accountability chart(Figures 4-5, 4-5A and 4-5B).

Hover ceiling – in ground effect charts (Figure4-6) and Hover ceiling – out of ground effectcharts (Figure 4-7) present hover performanceas allowable gross weight for conditions of HPand OAT. These hover ing weights areobtainable in zero wind conditions. Eachchart is divided into two areas: Area A (nonshaded area) and Area B (shaded area).

For the data presented below 14,000 ft HD,Area A of the hover ceiling charts presentshover performance (relative to GW) forconditions where adequate control marginsexist for all relative wind conditions up to 35knots for lateral CG not exceeding ±2.5 inches(±63 mm); and up to 17 knots, for lateral CGnot exceeding ±4.0 inches (±102 mm); for

σ

σ

σ

σ

σ


17 DEC 2002 4-5

hover, takeoff and landing. Area B of thehover ce i l ing char ts presents hoverperformance (relative to GW) for conditionswhere adequate control margins exist forrelative winds within ±45° of the nose ofhelicopter up to 35 knots for lateral CG notexceeding ±2.5 inches (±63 mm), and up to 17knots for lateral CG not exceeding ±4.0 inches(±102 mm); for hover, takeoff and landing.

For data presented between 14,000 and 17,000ft HD, Area A of the hover ceiling chartspresents hover performance (relative to GW)for conditions where adequate controlmargins exist for all relative wind conditionsup to 20 knots for lateral CG not exceeding±2.5 inches (±63 mm); for hover, takeoff andlanding. Area B of the hover ceiling chartspresents recommended azimuth for takeoffand landing for all relative winds within ±30°of nose of helicopter for lateral CG notexceeding ±2.5 inches (±63 mm).

For data presented above 17,000 ft HD, thereis no Area A. Area B presents hoverperformance (relative to GW) for conditionswhere adequate control margins exist for allrelative winds within ±30° of the nose of thehelicopter for lateral CG not exceeding ±2.5inches (±63 mm); for hover, takeoff andlanding

The following example uses a Hover ceilingchart at takeoff power. The example is typicalfor use with all other Hover ceiling charts.

EXAMPLE:

What IGE GW hover capability could beexpected for the following conditions:

A. HEATER and ANTI ICE – OFFB. HP – 6000 feetC. OAT – +20°CD. TAKE OFF POWER

SOLUTION:

Use Hover ceiling IGE – takeoff power chart(sheet 1 of Figure 4-6).

A. Enter OAT scale at +20 °C.B. Move upward to 6000 feet HP curve.C. Move horizontally to +20 °C curve.D. Drop down to read maximum

external gross weight of 5400 pounds (IGE hover capability exceeds maximum internal GW of 5000 pounds).

4-6. NOT USED

4-7. CLIMB AND DESCENT

4-7-A. CLIMB

Rate of climb charts are presented for variouscombinations of power settings and ENGINEANTI-ICING switch positions. Refer to Figures4-8 and 4-9.

Recommended best rate of climb airspeed is60 KIAS.

Reduce rate of climb data 100 feet per minutewhen operating with any combination ofdoor(s) removed.

The following example uses a Rate of climbchart at takeoff power. The example is typicalfor use with all other Rate of climb charts.

EXAMPLE:

Find the maximum rate of climb that can beattained using takeoff power under thefollowing conditions:

HEATER OFF

ENGINE ANTI-ICING OFF

OAT 10°C

HP 14,000 feet

GW 3500 pounds


4-6 17 DEC 2002

SOLUTION:

Enter appropriate gross weight chart (sheet 3of Figure 4-8). At HP scale of 14000 feetproceed horizontally to temperature of 10°C.Drop down vertically and read a rate of climbof 1700 feet per minute.

4-7-B. AUTOROTATION

Refer to Figure 4-10 for autorotational glidedistance as a function of altitude.

4-8. AIRSPEED CALIBRATION

Refer to Figure 4-11 for airspeed installationcorrection during level flight and climb.

4-9. NOT USED

4-10. NOISE LEVELS

4-10-A. FAR PART 36 STAGE 2NOISE LEVEL

This aircraft is certified as a Stage 2 helicopteras prescribed in FAR Part 36, Subpart H, forgross weights up to and including thecertificated maximum takeoff and landingweight of 5000 pounds (2268 kilograms).There are no operating limitations to meet anyof the noise requirements.

The following noise level complies with FARPart 36, Appendix J, Stage 2 noise levelrequirements. It was obtained by analysis ofapproved data from noise tests conductedunder the prov is ions of FAR Par t 36 ,Amendment 36-20.

The certified flyover noise level for the Model407 is 85.1 dBA SEL.

NOTE

No determination has been made bythe certifying authorities that the noise

levels of this aircraft are or should beacceptab le o r unacceptab le foroperations at, into, or out of anyairport.

VH is defined as the airspeed in level flightobtained using the minimum specificationengine torque corresponding to maximumcontinuous power available for sea level, 25°C(77°F) ambient conditions at the relevantmaximum certificated weight. The value of VHthus defined for this aircraft is 127 KTAS.

4-10-B. CANADIAN AIRWORTHINESS MANUAL CHAPTER 516 AND ICAO ANNEX 16 NOISE LEVEL

This aircraft complies with the noise emissionstandards applicable to the aircraft as set outby the In te rna t iona l C iv i l Av ia t ionOrganization (ICAO) in Annex 16, Volume 1,Chapter 11, for gross weights up to andincluding the certificated maximum takeoffand landing weight of 5000 pounds (2268kilograms). There are no operating limitationsto meet any of the noise requirements.

The following noise level complies with ICAOAnnex 16, Volume 1, Chapter 11 noise levelrequirements. It was obtained by analysis ofapproved data from noise tests conductedunder the provisions of ICAO Annex 16,Volume 1, Third Edition-1993.

The flyover noise level for the Model 407 is84.6 dBA SEL.

NOTE

ICAO Annex 16, Volume 1, Chapter 11approval is applicable only afterendorsement by the Civil AviationAuthority of the country of aircraftregistration.


17 DEC 2002 4-7

4

Figure 4-1. Power assurance check


4-8 17 DEC 2002

Figure 4-2. Density altitude

-50 -40 -30 -20 -10 0 10 20 30 40 50 60OUTSIDE AIR TEMPERATURE - °C

-8

-6

-4

-2

0

2

4

6

8

10

12

14

16

18

20

22

24

DENS

ITY

ALTI

TUDE

- FT

x 1

000

0.90

0.92

0.94

0.96

0.98

1.00

1.02

1.04

1.06

1.08

1.10

1.12

1.141.16

1.181.20

1.221.241.261.281.301.321.341.361.381.401.421.441.46

25,000

SEA LEVEL

STANDARD DAY

5,000

-5,000

10,000

15,000

20,000

PRESSURE ALTITUDE - FT

-10,000

1 /

σ

EXAMPLE: If OAT is -15°C andHP is 7000 ft, HD is 5000 ft

and is 1.081 / σ

M407_FM-1__FIG_4-2.WMF


17 DEC 2002 4-9

Figure 4-3. Altitude vs gross weight for height – velocity diagram

ALTITUDE VS GROSS WEIGHTFOR HEIGHT-VELOCITY DIAGRAM

13 14 15 16 17 18 19 20 21 22

GROSS WEIGHT - KG x 100

-40 -20 0 20 40 60

OAT - °C

30 34 38 42 46 50

GROSS WEIGHT - LBS x 100

8000

6000

4000

2000

SEA LEVEL

Hp - F

T10

,000

14,000 FT HD

MA

XIM

UM

OA

T

-200

0

MIN

IMU

M O

AT

LIM

IT

12,0

00

14,0

00

MA

XIM

UM

INT

ER

NA

L G

RO

SS

WE

IGH

T

DEMONSTRATEDTO 9000 FEET

DENSITY ALTITUDE

9000 FT HD

GR

OSS W

EIGH

T REG

ION

S

A

B

C

D



4-10 17 DEC 2002

Figure 4-4. Height – velocity diagram

0 10 20 30 40 50 60 70 80 90 100 110 120

INDICATED AIRSPEED - KNOTS

0

50

100

150

200

250

300

350

400

450

500

550

600

650

700

750

800

SK

ID H

EIG

HT

- F

EE

T

HEIGHT-VELOCITY DIAGRAMFOR GROSS WEIGHT REGIONS A TO D

0

20

40

60

80

100

120

140

160

180

200

220

240

SK

ID H

EIG

HT

- M

ET

ER

S

NOTE: LOW HOVER POINT IS AT 6 FT SKID HEIGHT

AVOID

AV

OID

AR

EA

S

REG

ION "A" BELO

W TH

IS LINE

RE

GIO

N "B" B

ELO

W TH

IS LINE

RE

GIO

N "C

" BE

LOW

THIS

LINE

RE

GIO

N "D

" BE

LOW

THIS

LINE



17 DEC 2002 4-11

Figure 4-5. Hover ceiling wind accountability chart – below 14,000 feet HD

± 45°

ALLAZIMUTHS

SHADED AREA (AREA B)

(± 102 mm); FOR HOVER, TAKEOFF AND LANDING.

PRESENTS HOVER PERFORMANCE (RELATIVE TO GW) FOR CONDITIONS WHERE ADEQUATE CONTROL MARGINS EXIST FOR RELATIVE WINDS WITHIN ± 45° OF NOSE OF HELICOPTER UP TO 35 KNOTS FOR LATERAL CG NOT EXCEEDING ± 2.5 INCHES (± 63 mm), AND UP TO 17 KNOTS FOR LATERAL CG NOT EXCEEDING ± 4.0 INCHES

NON-SHADED AREA (AREA A)

(± 102 mm); FOR HOVER, TAKEOFF AND LANDING.

PRESENTS HOVER PERFORMANCE (RELATIVE TO GW) FOR CONDITIONS WH ERE A DEQUATE CO NTROL MARGINS EXIST FOR ALL RELATIVE WIND CONDITIONS UP TO 35 KNOTS FOR LATERAL CG NOT EXCEEDING ± 2.5 INCHES (± 63 mm), AND UP TO 17 KNOTS FOR LATERAL CG NOT EXCEEDING ± 4.0 INCHES

ALTITUDE BELOW14000 FT HD

M407_FM-1_FIG_4-5.WMF


4-12 17 DEC 2002

Figure 4-5A. Hover ceiling wind accountability chart – between 14,000 and 17,000 feet HD

ALLAZIMUTHS

SHADED AREA (AREA B)PRESENTS RECOMMENDED AZIMUTH FOR TAKEOFF AND LANDING FOR ALL RELATIVE WINDS WITHIN ± 30° OF NOSE OF HELICOPTER FOR LATERAL CG NOT EXCEEDING ± 2.5 INCHES (± 63 mm).

NON-SHADED AREA (AREA A)PRESENTS HOVER PERFORMANCE (RELATIVE TO GW) FOR CONDITIONS W HERE ADEQUATE CONTROL MARGINS EXIST FOR ALL RELATIVE WIND CONDITIONS UP TO 20 KNOTS FOR LATERAL CG NOT EXCEEDING ± 2.5 INCHES (± 63 mm); FOR HOVER, TAKEOFF AND LANDING.

ALTITUDE BETWEEN14000 AND 17000 FT HD

± 30°

M407_FM-1__FIG_4-5A.WMF


17 DEC 2002 4-13

Figure 4-5B. Hover ceiling wind accountability chart – above 17,000 feet HD

ALTITUDE ABOVE17000 FT HD

SHADED AREA (AREA B)PRESENTS HOVER PERFORMANCE ( R E L A T I V E T O G W ) F O R CONDITIONS WHERE ADEQUATE CONTROL MARGINS EXIST FOR ALL WINDS DIRECTLY OFF THE NOSE OF THE HELICOPTER; FOR HOVER, TAKEOFF AND LANDING.

M407_FM-1_FIG_4-5B.WMF


4-14 17 DEC 2002

Figure 4-6. Hover ceiling IGE (sheet 1 of 8)

HOVER CEILINGIN GROUND EFFECT

TAKEOFF POWERENGINE RPM 100%GENERATOR 180 AMPS

SKID HEIGHT 4 FT (1.2 METER)HEATER AND ANTI-ICE OFF

BASIC INLET

NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART

14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26

GROSS WEIGHT - KG x 100M

AXIM

UM

OA

T

20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C51.7

5040

30

20

10

0

-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FM-1__FIG_4-6_(1_OF_8).WMF

MAX DEMONSTRATED HD


17 DEC 2002 4-15




SKID HEIGHT 4 FT (1.2 METER)ANTI-ICE ONBASIC INLET


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

5

0-10

-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FM-1__FIG_4-6_(2_OF_8).WMF

MAX DEMONSTRATED HD


4-16 17 DEC 2002




SKID HEIGHT 4 FT (1.2 METER)HEATER ON

BASIC INLET


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

2010

0-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FM-1__FIG_4-6_(3_OF_8).WMF

MAX DEMONSTRATED HD


17 DEC 2002 4-17




SKID HEIGHT 4 FT (1.2 METER)HEATER AND ANTI-ICE ON

BASIC INLET


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FM-1__FIG_4-6_(4_OF_8).WMF

MAX DEMONSTRATED HD


4-18 17 DEC 2002



MAXIMUM CONTINUOUS POWERENGINE RPM 100%GENERATOR 180 AMPS


BASIC INLET


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


MA

XIM

UM

OA

T20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

OAT - °C

51.750

4030

20

100

-10-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

EX

TE

RN

AL

= 6

000

LB

MA

X G

W IN

TE

RN

AL

= 50

00 L

B

M407_FM-1__FIG_4-6_(5_OF_8).EMF

MAX DEMONSTRATED HD


17 DEC 2002 4-19






14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FM-1__FIG_4-6_(6_OF_8).WMF

MAX DEMONSTRATED HD


4-20 17 DEC 2002





BASIC INLET


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

2010

0-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FM-1__FIG_4-6_(7_OF_8).EMF

MAX DEMONSTRATED HD


17 DEC 2002 4-21





BASIC INLET


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FM-1__FIG_4-6_(8_OF_8).WMF

MAX DEMONSTRATED HD


4-22 17 DEC 2002

Figure 4-7. Hover ceiling OGE (sheet 1 of 8)

HOVER CEILINGOUT OF GROUND EFFECT



BASIC INLET


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

MA

XIMU

M O

AT

OAT - °C

51.750

40

30

20

10

0

-10

-30

-40

-20

M407_FM-1__FIG_4-7_(1_OF_8).WMF

MAX DEMONSTRATED HD


17 DEC 2002 4-23






14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

OAT - °C

-20

50

-10

-30

-40

M407_FM-1__FIG_4-7_(2_OF_8).WMF

MAX DEMONSTRATED HD


4-24 17 DEC 2002





BASIC INLET


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

OAT - °C

-2020

100

-10

-30-40

M407_FM-1__FIG_4-7_(3_OF_8).WMF

MAX DEMONSTRATED HD


17 DEC 2002 4-25





BASIC INLET


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

OAT - °C

-20

50

-10

-30

-40

M407_FM-1__FIG_4-7_(4_OF_8).WMF

MAX DEMONSTRATED HD


4-26 17 DEC 2002





BASIC INLET


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA BM

AX

IMU

M O

AT

OAT - °C

51.750

4030

200

-10-30

-4010

-20

M407_FM-1__FIG_4-7_(5_OF_8).EMF

MAX DEMONSTRATED HD


17 DEC 2002 4-27






14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

OAT - °C

-205

0-10

-30-40

M407_FM-1__FIG_4-7_(6_OF_8).WMF

MAX DEMONSTRATED HD


4-28 17 DEC 2002





BASIC INLET


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

OAT - °C

-2020

100

-10

-30-40

M407_FM-1__FIG_4-7_(7_OF_8).EMF

MAX DEMONSTRATED HD


17 DEC 2002 4-29





BASIC INLET


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

OAT - °C

-205

0-10

-30-40

M407_FM-1__FIG_4-7_(8_OF_8).EMF

MAX DEMONSTRATED HD


4-30 17 DEC 2002

Figure 4-8. Rate of climb – takeoff power (sheet 1 of 10)

RATE OF CLIMB

REDUCE RATE OF CLIMB 225 FT/MIN ABOVE15,000 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)

Gross Weight 3000 lb (1361 kg)


60 KIASHEATER OFFBASIC INLET

0 200 400 600 800 1000 1200 1400 1600 1800 2000

RATE OF CLIMB - FT/MIN

0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10

RATE OF CLIMB - M/SEC

0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MAX O

AT LIM

IT

OAT-°C

30

20

10

0

-10

M407_FM-1__FIG_4-8_(1_OF_10).WMF


17 DEC 2002 4-31


RATE OF CLIMB




60 KIASHEATER ON

BASIC INLET

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

-20

20

10

0

-10

-30

M407_FM-1__FIG_4-8_(2_OF_10).WMF


4-32 17 DEC 2002


RATE OF CLIMB





0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

-20-30-40

40

M407_FM-1__FIG_4-8_(3_OF_10).WMF


17 DEC 2002 4-33


RATE OF CLIMB

REDUCE RATE OF CLIMB 190 FT/MIN ABOVE7500 FT Hp FOR ANTI-ICE ON (5°C AND COLDER)



60 KIASHEATER ON

BASIC INLET

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OAT LIM

ITOAT-°C

20

10

0

-10

-20

-30-40

M407_FM-1__FIG_4-8_(4_OF_10).WMF


4-34 17 DEC 2002


RATE OF CLIMB





0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

-20-30-40

40

50

M407_FM-1__FIG_4-8_(5_OF_10).WMF


17 DEC 2002 4-35


RATE OF CLIMB




60 KIASHEATER ON

BASIC INLET

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MAX OAT LIMIT OAT-°C

20

10

0

-10

-20

-30

-40

M407_FM-1__FIG_4-8_(6_OF_10).WMF


4-36 17 DEC 2002


RATE OF CLIMB





0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

-20

-30

-40

40

50

M407_FM-1__FIG_4-8_(7_OF_10).WMF


17 DEC 2002 4-37


RATE OF CLIMB




60 KIASHEATER ON

BASIC INLET

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

20

10

0

-10

-20 -30

-40

M407_FM-1__FIG_4-8_(8_OF_10).WMF


4-38 17 DEC 2002


RATE OF CLIMB





0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

-20

-30

-40

40

50

M407_FM-1__FIG_4-8_(9_OF_10).WMF


17 DEC 2002 4-39


RATE OF CLIMB




60 KIASHEATER ON

BASIC INLET

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 10

00

ME

TE

RS

OAT-°C

20

10

0

-10

-20

-30 -40

M407_FM-1__FIG_4-8_(10_OF_10).WMF


4-40 17 DEC 2002

Figure 4-9. Rate of climb – maximum continuous power (sheet 1 of 10)

RATE OF CLIMB





0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 10

00

ME

TE

RS

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

40

50

-20

-30

M407_FM-1__FIG_4-9_(1_OF_10).WMF


17 DEC 2002 4-41


RATE OF CLIMB




60 KIASHEATER ON

BASIC INLET

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 10

00

ME

TE

RS

MA

X O

AT LIM

ITOAT-°C

20

10

0

-10

-20

-30

-40

M407_FM-1__FIG_4-9_(2_OF_10).WMF


4-42 17 DEC 2002


RATE OF CLIMB





0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 10

00

ME

TE

RS

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

-20

-30

-40

40

50

M407_FM-1__FIG_4-9_(3_OF_10).WMF


17 DEC 2002 4-43


RATE OF CLIMB




60 KIASHEATER ON

BASIC INLET

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 10

00

ME

TE

RS

MA

X OA

T LIMIT

OAT-°C

20

10

0

-10

-20

-30

-40

M407_FM-1__FIG_4-9_(4_OF_10).WMF


4-44 17 DEC 2002


RATE OF CLIMB





0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

-20

-30

-40

40

50

M407_FM-1__FIG_4-9_(5_OF_10).WMF


17 DEC 2002 4-45


RATE OF CLIMB

REDUCE RATE OF CLIMB 170 FT/MINFOR ANTI-ICE ON (5°C AND COLDER)



60 KIASHEATER ON

BASIC INLET

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

20

10

0

-10

-20

-30

-40

M407_FM-1__FIG_4-9_(6_OF_10).WMF


4-46 17 DEC 2002


RATE OF CLIMB





0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

-20

-30

-40

40

50

0-20

-40

M407_FM-1__FIG_4-9_(7_OF_10).WMF


17 DEC 2002 4-47


RATE OF CLIMB




60 KIASHEATER ON

BASIC INLET

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

20

10

0

-10

-20

-30

-40

-20-4

0

M407_FM-1__FIG_4-9_(8_OF_10).WMF


4-48 17 DEC 2002


RATE OF CLIMB





0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

-20

-30

-40

40

50

M407_FM-1__FIG_4-9_(9_OF_10).WMF


17 DEC 2002 4-49


RATE OF CLIMB




60 KIASHEATER ON

BASIC INLET

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

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M407_FM-1__FIG_4-9_(10_OF_10).WMF


4-50 17 DEC 2002

Figure 4-10. Autorotation glide distance

AUTOROTATION GLIDE DISTANCE

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

Distance Over Ground (Nautical Miles)

0

2000

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NOTE: AUTOROTATIONAL DESCENTPERFORMANCE IS A FUNCTION OFAIRSPEED AND IS ESSENTIALLYUNAFFECTED BY DENSITY ALTITUDEAND GROSS WEIGHT.

Minimum Rate ofDescent Airspeed 55 KIAS

Maximum Glide DistanceAirspeed 80 KIAS

M407_FM-1__FIG_4-10.EMF


17 DEC 2002 4-51/4-52

Figure 4-11. Airspeed installation correction

AIRSPEED INSTALLATION CORRECTION TABLEKCAS = (KIAS – INSTRUMENT ERROR – POSITION ERROR)

NOTE: This chart assumes zero instrument error.

KIAS CLIMBKCAS

LEVEL FLIGHTKCAS

20 – – 22

30 30 33

40 37 43

50 47 52

60 58 63

70 69 73

80 78 82

90 87 92

100 95 100

110 – – 110

120 – – 121

130 – – 131

140 – – 144

M407_FM-1_FIG_4-11.WMF

BHT-407-FM-1

17 DEC 2002 5-1

Section 5WEIGHT AND BALANCE

5x


Subject

TABLE OF CONTENTS

INTRODUCTION ......................................................................... 5-1.................... 5-3EMPTY WEIGHT CENTER OF GRAVITY .................................. 5-2.................... 5-3

EMPTY WEIGHT ................................................................. 5-2-A................ 5-3CENTER OF GRAVITY ....................................................... 5-2-B................ 5-3

GROSS WEIGHT CENTER OF GRAVITY.................................. 5-3.................... 5-3USEFUL LOADS ................................................................. 5-3-A................ 5-3CENTER OF GRAVITY ....................................................... 5-3-B................ 5-3

DOORS OPEN OR REMOVED................................................... 5-4.................... 5-4DOOR WEIGHTS AND MOMENTS .................................... 5-4-A................ 5-4BALLAST ADJUSTMENT................................................... 5-4-B................ 5-4

COCKPIT AND CABIN LOADING .............................................. 5-5.................... 5-4LONGITUDINAL LOADING ................................................ 5-5-A................ 5-5MOST FORWARD AND MOST AFT CG ............................ 5-5-B................ 5-5ALTERNATE LOADING...................................................... 5-5-C................ 5-5CABIN FLOOR LOADING................................................... 5-5-D................ 5-5

BAGGAGE COMPARTMENT LOADING ................................... 5-6.................... 5-5FUEL LOADING.......................................................................... 5-7.................... 5-6SAMPLE LOADING PROBLEM ................................................. 5-8.................... 5-6


Page

LIST OF FIGURES

Fuselage stations....................................................................... 5-1.................... 5-7Buttock lines .............................................................................. 5-2.................... 5-8Fuel center of gravity................................................................. 5-3.................... 5-9

Title Number

TableNumberPage

LIST OF TABLES

Door Weights and Moments (U.S.) ........................................... 5-1................... 5-10Cabin and baggage loading (U.S.)............................................ 5-2................... 5-11Cabin and baggage loading (Metric) ........................................ 5-3................... 5-12Fuel Loading (U.S.) .................................................................... 5-4................... 5-13Fuel Loading (Metric)................................................................. 5-5................... 5-14

BHT-407-FM-1

5-2 17 DEC 2002

Title Number

TableNumberPage

LIST OF TABLES (CONT)

Fuel Density vs Temperature.................................................... 5-6 .................. 5-15Sample Loading Problem (U.S.) ............................................... 5-7 .................. 5-16Sample Loading Problem (Metric)............................................ 5-8 .................. 5-17Weight and Balance Worksheet (U.S.)..................................... 5-9 .................. 5-18Weight and Balance Worksheet (Metric) ................................. 5-10 ................ 5-19

BHT-407-FM-1

17 DEC 2002 5-3


55-1. INTRODUCTION

This section presents loading information andinstructions necessary to ensure that flightcan be performed within approved grossweight and center of gravity limitations asdefined in Section 1.

5-2. EMPTY WEIGHT CENTER OF GRAVITY

5-2-A. EMPTY WEIGHT

The empty weight condition consists of thebasic helicopter with required equipment,optional equipment kits, transmission andgearbox oils, hydraulic fluid, unusable fuel,undrainable engine oil, and fixed ballast. Theempty weight and center of gravity arerecorded on the Actual Weight Record, a copyof which should be carried in the helicopter toenable weight and balance computations.


An empty weight center of gravity chart isprovided in maintenance manual as a guide tosimplify computing ballast requirements. Thischar t was der ived f rom gross we ightlongitudinal center of gravity limits shown inSection 1, using most forward and most aftuseful loads for standard seating and fuel.

NOTE

Empty weight center of gravity chartis not va l id i f he l icop ter has anonstandard fuel system or seatingarrangement.

5-3. GROSS WEIGHT CENTER OF GRAVITY

Gross weight condition is empty weightcondition plus useful load.

5-3-A. USEFUL LOADS

Useful load consists of usable fuel, engine oil,crew, passengers, baggage and cargo.Combinations of these items which have mostadverse effect on helicopter center of gravityare known as most forward and most aftuseful loads. Whenever cargo and/or baggageare carr ied, these useful loads may bedifferent for each flight, and weight andbalance must be computed to ensure grossweight and center of gravity will remain withinlimits throughout flight.

Standard most forward and most aft usefulloads are combinations of fuel, crew andpassenger loading only. These loads, inconjunction with empty weight center ofgravity chart, allow passengers only (nobaggage or other cargo) to be carried withinappropriate weight l imitat ions withoutcomputing center of gravity for each flight.

If helicopter has a nonstandard fuel system orseating arrangement, or is not ballasted inaccordance with empty weight center ofgravity chart in maintenance manual, pilotmust determine weight and balance to ensuregross weight and center of gravity will remainwithin limits throughout each flight.


It is the responsibility of the pilot to ensurethat helicopter is properly loaded to maintain

BHT-407-FM-1

5-4 17 DEC 2002

center of gravity throughout each flight withingross weight center of gravity limits shown inSection 1 or appropriate supplement. Grossweight longitudinal and lateral center ofgravity can be calculated using Actual WeightRecord, diagrams and loading tables in thissection and loading tables in applicable flightmanual supplements.

When carry ing baggage, cargo o rnonstandard loads , e f fec ts o f fue lconsumption and addi t ion/delet ion ofpassengers , baggage or cargo a tintermediate points should be checked priorto flight.

Significant fuselage stations and buttocklines are shown in Figures 5-1 and 5-2 to aidin weight and balance computations.

5-4. DOORS OPEN OR REMOVED

When one or more cabin doors are removed,helicopter may exceed gross weight center ofgravity limits during flight. If using Weightempty center of gravity chart, refer to BHT-407-MM-1, a ballast adjustment to offsetmoment change is necessary (Table 5-1).Otherwise, gross weight center of gravityshould be computed for each flight.

5-4-A. DOOR WEIGHTS AND MOMENTS

Following table presents weight and momentadjustments for cabin doors. Sign conventionfor buttock lines used to compute lateralmoments are:

1. Left is negative.

2. Right is positive.

Example:

When removing a left door only, subtractpositive weight value and negative momentvalue shown in table. Net effect on helicopteris a reduction in weight and a shift in lateralCG to right (positive direction).

5-4-B. BALLAST ADJUSTMENT

Following check can be made to determine ifa ballast adjustment is necessary after doorsare removed or installed.

1. For helicopters without ballast orwith nose ballast, apply weight andmoment changes to most aft usefulload condition to determine if anincrease in nose ballast is required,or a reduction is allowed.

2. For helicopters with tail ballast,apply weight and moment changesto most fo rward use fu l loadcondition to determine if a reductionin tai l bal last is a llowed, or anincrease is required.

NOTE

Ballast changes are performed bymaintenance personnel. After anyballast change, Actual Weight Recordmust be revised to show new weightempty condition.

5-5. COCKPIT AND CABIN LOADING

Loading tables (Tables 5-2 and 5-3) provideweights and moments for each passengerlocat ion , l i t te r pat ient and baggagecompartment in both U.S. and metric units.

To find moments for weights in excess ofthose shown on tables, multiply weight byfuselage station at which center of gravity of

ACTION MOMENT CHANGELEFT DOOR RIGHT DOOR

Remove Positive (+) Negative (-)Install Negative (-) Positive (+)

BHT-407-FM-1

17 DEC 2002 5-5

the object is located. An alternate method isto calculate amount of weight in excess ofmaximum weight listed on table, then readmoment for this excess weight from table andadd it to moment for maximum weight shownon table. This will give desired moment for theobject.

5-5-A. LONGITUDINAL LOADING

1. A minimum weight of 170 pounds(77.1 k i lograms) is required incockpit at fuselage station 65.0when the empty weight center ofgravity chart is used.

2. Passenger seating is unrestricted.

3. Cargo loading is restricted only byfloor load limit. Refer to Section 1.

5-5-B. MOST FORWARD AND MOST AFT CG

When using empty weight center of gravitychart, following combinations of crew, fueland passenger loading wil l have mostextreme effects on longitudinal center ofgravity, assuming standard weights for allcrew and passengers.

1. Most forward CG will occur withforward and mid seats occupied andfuel quantity of 74.8 gallons (283.0liters).

2. Most aft CG will occur with oneforward seat occupied (pilot) andfuel quantity of 28.4 gallons (107.5liters).

Since center of gravity of aft passengers is onaft limit, weight of passengers is not includedin most aft useful load. However when mostaft center of gravity of a configuration isforward of aft limit, addition of aft passengerswill shift center of gravity further aft, andshould be included in computation.

5-5-C. ALTERNATE LOADING

Gross weight center of gravity chart must beused to determine cabin loading requirementsunder following conditions:

1. Whenever cargo and/or baggage arecarried.

2. When actual passenger weights areused.

3. When seating arrangement and/orfuel system are non-standard.

4. When performing specialty missions,such as hoisting or rappelling.

5-5-D. CABIN FLOOR LOADING

Cabin floor is structurally designed for 75pounds per square foot (3.7 kilograms per 100square centimeters).

5-6. BAGGAGE COMPARTMENT LOADING

When we ight is loaded in to baggagecompartment, the pilot is required to computeweight and balance, regardless of passengerloading.

Baggage compartment is s tructura l lydesigned for 86 pounds per square foot (4.2kilograms per 100 square centimeters) for atotal weight of 250 pounds (113.4 kilograms).

Loading of baggage compartment should befrom front to rear. Load shall be secured totiedown fittings if shifting of load in flightcould result in structural damage to baggagecompartment or in gross weight center ofgravity being exceeded.

If load is not secured, center of gravity mustbe computed with load in most adverseposition.

BHT-407-FM-1

5-6 17 DEC 2002

5-7. FUEL LOADING

Longitudinal center of gravity of fuel shifts asit is consumed (Figure 5-3). Extreme effects offuel consumption on helicopter center ofgravity for standard fuel system are asfollows:

1. Critical fuel for computing mostforward useful load is 74.8 gallons(283.0 liters).

2. Critical fuel for computing most aftuseful load is 28.4 gallons (107.5liters).

Fuel loading tables (Tables 5-4 and 5-5) listusable fuel quantities, weight and moments inboth U.S. and metric units.

Fuel density vs temperature (Table 5-6), isprovided to calculate fuel weight variation forequivalent volumes of fuel caused by achange in temperature. For example weight of

127.8 gallons (full fuel) of JP-5 at -40°F is913.8 pounds (414.5 kilograms) versus 869.0pounds (394.1 kilograms) shown on Fuelloading chart (Tables 5-4 and 5-5).

5-8. SAMPLE LOADING PROBLEM

A sample loading problem showing derivationof critical gross weights and center of gravitylocations for a typical mission is presented inU.S. and metric units (Tables 5-7 and 5-8).Method shown derives a gross weight withzero fuel for each load condition to bechecked, then adds appropriate fuel weightand moment read directly from Fuel loadingtable. Center of gravity for each condition iscalculated by dividing total moment by totalweight.

Forms have been provided (Tables 5-9 and 5-10) in both U.S. and Metric Units, to aid incomputing critical load conditions for a flight.

BHT-407-FM-1

17 DEC 2002 5-7

Figure 5-1. Fuselage stations

BHT-407-FM-1

5-8 17 DEC 2002

Figure 5-2. Buttock lines

BHT-407-FM-1

17 DEC 2002 5-9

Figure 5-3. Fuel center of gravity

BHT-407-FM-1

5-10 17 DEC 2002

Table 5-1. Door Weights and Moments (U.S.)

BHT-407-FM-1

17 DEC 2002 5-11

Table 5-2. Cabin and baggage loading (U.S.)

BHT-407-FM-1

5-12 17 DEC 2002

Table 5-3. Cabin and baggage loading (Metric)

BHT-407-FM-1

17 DEC 2002 5-13

Table 5-4. Fuel Loading (U.S.)

BHT-407-FM-1

5-14 17 DEC 2002

Table 5-5. Fuel Loading (Metric)

BHT-407-FM-1

17 DEC 2002 5-15

Table 5-6. Fuel Density vs Temperature

BHT-407-FM-1

5-16 17 DEC 2002

Table 5-7. Sample Loading Problem (U.S.)

BHT-407-FM-1

17 DEC 2002 5-17

Table 5-8. Sample Loading Problem (Metric)

BHT-407-FM-1

5-18 17 DEC 2002

Table 5-9. Weight and Balance Worksheet (U.S.)

BHT-407-FM-1

Table 5-10. Weight and Balance Worksheet (Metric)

17 DEC 2002 5-19/5-20

BHT-407-FM-1

Appendix AOPTIONAL EQUIPMENT SUPPLEMENTS

ATABLE OF CONTENTS


Optional Equipment............................................................................... A-1 .......... A-3

LIST OF TABLES


Flight Manual Supplements for Optional Equipment ......................... A-1 ....... A-3

19 FEB 2007—Rev. 5———A-1/A-2

http://127.0.0.1:1584/407-FM-1 A.pdf

BHT-407-FM-1

19 FEB 2007—Rev. 5———A-3

Appendix AOPTIONAL EQUIPMENT SUPPLEMENTS

AA-1. OPTIONAL EQUIPMENT

Bell Helicopter Textron's policy is one ofcontinuous product improvement and Bellreserves the right to incorporate changes,make additions to and improve its productswithout imposing any obligation upon thecompany to furnish for or instal l such

changes, additions, improvement, etc., on itsproducts previously manufactured.

The following items may be installed on thebasic helicopter by authorized personnel.Only the opt ional equipment l is ted inTable A-1 require a Flight Manual Supplement.

Table A-1: Flight Manual Supplements for Optional Equipment

NAME OF EQUIPMENT KIT NUMBERCERTIFIED

DATEREVISIONCURRENT

BHT-407-FMS-1 Lightweight Emergency Flotation Landing Gear

407-706-008 11 APR 96 Original

BHT-407-FMS-2 High Skid Gear

407-706-007 14 FEB 96 Original

BHT-407-FMS-3 Particle Separator

206-706-212 1 MAR 96 Reissue16 DEC 02

BHT-407-FMS-4 Snow Deflector

206-706-208 1 MAR 96 Reissue16 DEC 02

BHT-407-FMS-5 Cargo Hook

206-706-341 14 FEB 96 Rev. 14 SEP 98

BHT-407-FMS-6 Auxiliary Fuel Kit

407-706-011 20 MAR 96 Original

BHT-407-FMS-7 Litter(s)

407-706-631or

407-799-100or

407-799-001

14 FEB 96 Rev. 116 SEP 99

BHT-407-FMS-8 Reserved


BHT-407-FMS-10 Helicopters Registered in U.S.A.

Canceled

BHT-407-FM-1

A-4———Rev. 5—19 FEB 2007







BHT-407-FMS-17 Cargo Tie-down Provisions Kit

407-705-201 1 APR 96 Original



BHT-407-FMS-20 KLN 89B GPS Navigator

407-705-001 14 FEB 96 Rev. 126 NOV 96

BHT-407-FMS-21 Fire Detection System

407-799-004or

407-706-015or

407-706-025

2 MAY 96 Reissue7 JUL 04

BHT-407-FMS-22 Auxiliary Vertical Fin Strobe Lights

407-899-023 10 MAY 96 Original

BHT-407-FMS-23 Ryan Traffic Collision Avoidance Device

407-899-022 15 MAY 96 Original


BHT-407-FMS-25 Quiet Cruise Mode

407-706-016 8 MAY 98 Reissue17 DEC 02

BHT-407-FMS-26 Modified Hydromechanical Unit

Canceled

BHT-407-FMS-27 FADEC Software Version 5.201

Canceled

BHT-407-FMS-28 Increased Internal Gross Weight

407-706-020 16 MAR 99 Reissue16 DEC 02

BHT-407-FMS-29Airspeed Actuated Pedal Stop

Canceled




DATEREVISIONCURRENT

BHT-407-FM-1

BHT-407-FMS-31Increased APC Starter Generator Load

407-706-026 15 JUN 05 Original

BHT-407-FMS-32Hanger Bearing Vibration Monitor Kit(not distributed to all customers)

407-706-023 15 JUL 03 Original

BHT-407-FMS-CAA United KingdomRegistered Helicopters

08 JAN 02 Original

BHT-407-FMS-IAC ARInterstate Aviation Committee — Aviation Register Commonwealth of Independent States

407-706-021 20 MAY 99 Original



DATEREVISIONCURRENT

19 FEB 2007—Rev. 5———A-5/A-6

BHT-407-FM-I-ATO

11MODEL 407ROTORCRAFT

FLIGHT MANUALFOR REPUBLIC OF PHILIPPINES

REGISTERED HELICOPTERSRepublic of the Philippines

Department of Transportation and CommunicationsAIR TRANSPORTATION OFFICE

NAIA, 1300 Pasay City, M.M

VALIDATED AIRCRAFT TYPE CERTIFICATENO. H1TSN-001

Inspect And Evaluated By:

L. t&JERA, JR.in er III

Reviewed By:

JO9E'C. ROBLES, JR. 6ATURNINO B. DELAChief, Aircraft Engineering Section Chief, Aviation Safety

(RET.) DATE: November 12, 1996

THE AVLAT70N REGULATORY AUTHORITY FOR THIS FLIGHT MANUAL IS THEAIR TRANSPORTATION OFFICE OF THE RUPUBLIC OF THE PHILIPPINES.

R.P. REGISTERED HELICOPTERS ARE APPROVED BY ATO IN ACCORDANCEWITH THE PROVISIONS OF REPUBLIC ACT 776 SERIES OF 1952 AS AMENDED BYPRESIDENTIAL DECREE. NO. 844. 1278. 1462 AND EXECUTIVE ORDER NO. 546.


Bell HelicopterCOPYRIGHT NOTICE

COPYRIGHT 1996BELL @ HELICOPTER TEXTRON INC.AND BELL HELICOPTER TEXTRON,

A DIVISION OF TEXTRON CANADA LTD.ALL RIGHTS RESERVED

A Subsidiary of Textron Inc


12 NOVEMBER 1996

BHT 407-FM-I-ATO

NOTICE PAGE

The pages contained herein shall be attached to the basic flight Manualwhen the helicopter is registered in the Republic of the Philippines.


Bell Helicopter Textron Inc.P.O. Box 482


NP

BHT-407-FM-I-ATO

LOG OF REVISIONS

Original .......... 0 ... November 12, 1996

PAGE

LOG OF PAGES

REVISION REVISIONNO. PAGE NO.

Title - NP ............................ 0A/B ................................... 0

NOTE

Revised text is indicated by a black vertical line. Insert latest revision pages: dispose ofsuperseded pages

A/B

Tom

m 0

BHT-407-FM-CTA

MODEL 9mroROTORCRAFT

FLIGHT MANUAL4, -' -

REGISTRATION NO.

DOCUMENT NO.

DIRECTOR AIRWORTHINESS BRANCHDEPARTMENT OF TRANSPORT

SERIAL NO.

FOREIGN AUNWR" REPRESENTATIVE

THIS MANOL AND ASSOCIATED SUPPLEMENTS ARE APPROVED INACCORDANCE WITH SECTION 21.29 OF RBHA 21 FOR BRAZILIAN REGISTEREDAIRCRAFT AND IS APPROVED BY THE D.O.T. ON BEHALF OF THE CENTROTECNICO AEROESPACIAL, WHEN HELICOPTER IS REGISTERED IN BRAZIL.

THIS AIRCRAFT SHOULD BE OPERATED IN ACCORDANCE WITH THELIMITATIONS AND INSTRUCTIONS HEREIN ESTABLISHED.


Bell HelicopterCOPYRIGHT NOTICE

COPYRIGHT 1998BELL 0 HELICOPTER TEXTRON INC.AND BELL HELICOPTER TEXTRON INC.A DIVISION OF TEXTRON CANADA LTD.

ALL RIGHTS RESERVED



CTA APPROVED - 3 APRIL 1998

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BHT 407-FM-CTA

NOTICE PAGE

The following Warning is not applicable to helicopters on which all kits andcustomizing installations have been qualified and approved by Bell Helicopter.

WARNING

THE HELICOPTER MAY CONTAIN INSTALLATIONS, PARTS, ORPROCESSES CERTIFIED BY PARTIES OTHER THAN BELLHELICOPTER TEXTRON. BELL HELICOPTER CAN NOTCONFIRM THAT SUCH INSTALLATIONS HAVE BEEN FULLYQUALIFIED OR CONFORMED TO BELL HELICOPTER DESIGNCRITERIA. AS A RESULT OF SUCH INSTALLATIONS, BELLHELICOPTER SUPPLIED DATA MAY NOT BE VALIDCONCERNING IN-FLIGHT HANDLING QUALITIES, WEIGHT ANDBALANCE, OR HELICOPTER PERFORMANCE. IF MULTIPLESTC KITS OR SIMILAR INSTALLATIONS ARE INCORPORATED,THERE MAY BE NO VALID TEST DATA TO QUALIFY THEHELICOPTER AS MODIFIED BY THESE INSTALLATIONS. FORREVISED DATA, CONTACT THE OWNER OF THE INSTALLEDSTC OR THE SUPPLIER FOR THE APPLICABLE APPROVAL OFEACH INSTALLATION.




NP

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BHT-407-FM-IAC AR

MODELr-1r


DOCUMENT N0.

REGISTRATION N0. SERIAL NO.

DATE L1 . l f

DIRECTOR - AI AFT CERTIFICATION BRANCHDEPARTM T OF TRANSPORT, CANADA

THIS MANUAL AND ASSOCIATED SUPPLEMENTS ARE APPROVED !NACCORDANCE WITH SECTIONS 2.10 AND 4.7 OF AP 21 FOR C1S REGISTEREDAIRCRAFT AND IS APPROVED BY THE D.O.T. ON BEHALF OF THE AVIATIONREGISTER OF THE INTERSTATE AVIATION COMMITTEE.

THIS AIRCRAFT SHOULD BE OPERATED !N ACCORDANCE WITH THELIMITATIONS AND INSTRUCTIONS HEREIN ESTABLISHED.


COPYRIGHT NOTICE

COPYRIGHT 1999BELL 0 HELICOPTER TEXTRON INC.AND BELL HELICOPTER TEXTRON INC.A DIVISION OF TEXTRON CANADA LTD.ALL RIGHTS RESERVED

Bell HelicopterA Subsidiary of Textron Inc.

POST OFFICE 80X 482 FORT WORTH, TEXAS 76101

20 MAY 1999

W-0

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(CN

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BHT-407-FM-IAC AR

NOTICE PAGE

The following Warning is not applicable to helicopters on which all kits andcustomizing installations have been qualified and approved by Bell Helicopter.

WARNING

THE HELICOPTER MAY CONTAIN INSTALLATIONS, PARTS,OR PROCESSES CERTIFIED BY PARTIES OTHER THANBELL HELICOPTER TEXTRON. BELL HELICOPTER CANNOT CONFIRM THAT SUCH ITEMS HAVE BEEN FULLYQUALIFIED OR CONFORMED TO BELL HELICOPTERDESIGN CRITERIA. BELL HELICOPTER SUPPLIED DATAMAY NOT BE VALID CONCERNING IN-FLIGHT HANDLINGQUALITIES, WEIGHT AND BALANCE, OR SIMILARINSTALLATIONS. THERE MAY BE NO VALID TEST DATA TOQUALIFY THE HELICOPTER AS MODIFIED BY THESE ITEMS.FOR REVISED DATA, CONTACT THE OWNER OF THEINSTALLED STC (OR EQUIVALENT CERTIFICATION) OR THESUPPLIER FOR THE APPLICABLE APPROVAL OF EACHINSTALLATION.

PROPRIETARY RIGHTS NOTICE

Manufacturer's Data portion of this manual is proprietary to BellHelicopter Textron Inc. Disclosure, reproduction, or use of these data forany purpose other than helicopter operation is forbidden without priorwritten authorization from Bell Helicopter Textron Inc.




NP

B HT-407-FM S-i

'!407ROTORCRAFT

FLIGHT MANUAL

SUPPLEMENTLiGHTWEIGHT EMERGENCYFLOTATION LANDING GEAR

407-706-008

CERTIFIED11 APRIL 1996

This supplement shall be attached to Model 407Flight Manual when 407-706-008 LightweightEmergency Flotation Landing Gear kit has beeninstalled.

Information contained herein supplementsinformation of basic Flight Manual. ForLimitations, Procedures, and Performance Datanot contained in this supplement, consult basicFlight Manual.

_______________ Bell Helicopterti4it.1iCOPYRIGHT NOTICE A Subsidiary of Textron Inc.

COPYRIGHT 1996BELL HELICOPTER INC. ST OFFICE BOX 482 • FORT WORTH. TEXAS 78101

AND BELL HELICOPTER TEXTRON INC.A DIVISION OF TEXTRON CANADA LTD. 11 APRIL 1 99ALL RIGHTS RESERVED

BHT-407-FMS-1

NOTICE PAGE


Bell Helicopter Textron Inc.P. 0. Box 482


NP /

BHT-407-FMS-1

LOG OF REVISIONS

Original ........... 0 11 Apr 1996

LOG OF PAGES

REVISION REVISIONPAGE NO. PAGE NO.

FLIGHT MANUAL C/D 0i/li 0

Title—NP 0 1—4 0A—B 0 5/6 0

NOTE

Revised text is indicated by a black vertical line. Insert latest revision pages; dispose ofsuperseded pages.

A

BHT-407-FMS-1 DOT APPROVED

LOG OF APPROVED REVISIONS

Original .0 .11 Apr 1996

APPROVED:"'I0, DATE: /1 ,'9pe q"

PCHIEF, FLIGHT TESTFORDIRECTOR — AIRWORTHINESS BRANCHDEPARTMENT OF TRANSPORT

B

BHT-407-FMS-1


Original .0 .11 Apr 1996

C/D

GENERAL INFORMATION

BHT-407-FMS-1

Lightweight emergency flotation landinggear kit (407-706-008) will allow helicopterto land in water during an emergencysituation. Kit consists of six skid mountedpop-out float bags, an inflation systemwith electrically operated solenoid valves,and attaching hardware. Two pneumaticcharging bottles, located on underside ofhelicopter, are interconnected by apneumatic line that will cause chargingbottle valve to open in event that its

solenoid valve fails while floats are beinginflated. Each float assembly is equippedwith an inlet check valve, high pressurerelief valve which opens at 5.25 PSIG0.25 PSI and a finger operated manualstop-cocklinflation valve. A GEN FAILcaution light alerts pilot of generatorfailure and of battery power possibly beinginsufficient to inflate floats. Float inflationtime is approximately 5 seconds.

i/u

DOT APPROVED BHT-407-FMS-1

Lsection 1 1LIMITA TIONS

1-3. TYPES OF OPERATION Maximum allowable airspeed, floatsinflated, is 60 KIAS.

Emergency floats are installed forassistance during emergency ditching.

1-6. WEiGHT AND CENTER OFGRAVITY

Maximum autorotation airspeed, floatsinflated, is 60 KIAS.

1-8. ALTiTUDEMaximum inflation altitude is 5000 feet H.

1-9. MANEUVERINGActual weight change shall be determinedafter kit is installed and ballast readjusted,if necessary, to return empty weight CG towithin allowable limits. Refer to Center ofgravity vs weight empty chart in BHT-407-MM-i.

1-7. AIRSPEED

1-7-A. FLOATS STOWED

Floats stowed, covers installed — Same asbasic helicopter.

1-7-B. FLOATS INFLATED

1-9-B. CLIMB AND DESCENT

Maximum rate of climb with floats inflatedis 1000 feet per minute.

1-20. INSTRUMENTMARK1NGS AND PLACARDS

FLOATARMINGJNFLATLON

ABOVE 60 KIASPROHIBITED

Maximum arming/inflation airspeed is 60KIAS.

1



2-3. PREFLIGHT CHECK1. Floats — Stowed.

2. Nitrogen lines — Condition andsecurity.

3. Float covers — Clean andsecured.

4. Float inflation cylinders — Checkfor proper inflation pressure vstemperature and altitude. Refer toplacard on cylinders. Checkelectrical connectors for security.

2-4. INTERIOR ANDPRESTART CHECK

4. FLOAT TEST and FLOAT ARMlights — Press C/W LT TESTbutton.

5. FLOAT TEST button — Press andhold.

6. FLOAT INFLATE button — Press;check FLOAT TEST lightilluminates. Release button,check light extinguishes.

7. FLOAT TEST button — Release.

8. FLOAT ARM switch — Up, guardopen. Check FLOAT ARM lightilluminates, then switch down,guard closed. Check lightextinguishes.

2-4-A. PREFLIGHT FLOAT SYSTEMCHECK

2

1. BATT switch — BATT. With GENswitch OFF, verify GEN FAIL lightilluminates.

CAUTION;4,444444,4,,4444

IF GEN FAIL LIGHT DOES NOTILLUMINATE, MONITORVOLTMETER TO DETERMINEGENERATOR OPERATION. IFVOLTAGE DROPS BELOW 25VOLTS, PERFORM GEN FAILCORRECTIVE ACTION PER TABLE3-1.

2. FLOAT ARM switch — Down,guard closed.

3. FLOATS circuit breaker — Checkin.


2-9-A. OVER WATEROPERATIONS

1. FLOAT ARM switch — Up, guardopen.

2. FLOAT ARM light — illuminated.

CAUTION

DURING FLIGHT AT ALTITUDESABOVE 500 FEET AGL AND ATAIRSPEEDS OF 60 KIAS ANDABOVE, SYSTEM SHOULD BEDEACTIVATED BY PLACINGFLOAT ARM SWITCH TO DOWNPOSITION AND CLOSING GUARD.

3. Rearm system prior to landing.


2-9-B. OVER LAND OPERATiONS

FLOAT ARM switch — Down, guardclosed.

2-10. DESCENT ANDLANDING

WARNIN

IF CG IS AFT OF STATION 126,PRACTICE TOUCHDOWN

AUTOROTATIONS SHALL BEAVOIDED DUE TO NOSEDOWNPITCHING.

RUN-ON LANDINGS, ON OTHERTHAN A HARD FIRM SURFACE,SHOULD BE EXERCISED WITHCAUTION.

NOTE

Tail-low run-on landings shouldbe avoided to prevent nosedownpitching.


3-1. INTRODUCTIONTable 3-1 presents fault conditions andcorrective actions for cautions lights.

Table 3-1.

Generator notconnected toDC buss.

WARNING

IF GEN FAIL LIGHT ILLUMINATES,BATTERY POWER MAY NOT BESUFFICIENT TO INFLATEFLOATS. IF VOLTAGE DROPSBELOW 25 VOLTS, PERFORMGEN FAIL CORRECTIVE ACTIONPER TABLE 3-1.

Verify fault with AMP orVOLT gage. Over land: GENswitch — RESET, then ON. IfGEN FAIL light remainsilluminated or voltage dropsbelow 25 volts, switch —OFF. Land as soon aspractical.

3

GEN FAIL

PANEL FAULT CORRECTIVE ACTIONWORDING CONDITION


Table 3-1. (Cont)PANELWORDING

FAULTCONDITION

CORRECTIVE ACTION

Over water: GEN switch —RESET, then ON. If GEN FAILlight remains illuminated orvoltage drops below 25 volts,switch — OFF. Turn oft allnonessential electricalequipment to conservebattery power. Land as soonas practical.

3-15. EMERGENCY FLOATINFLATION

1. Reduce airspeed below maximuminflation airspeed — 60 KIAS.

2. Establish autorotation or lowpower descent at approximately500 feet per minute.

NOTE

4. FLOAT ARM light — illuminated.

CAUTION

MAXIMUM INFLATION ALTiTUDEIS 5000 H.

5. FLOAT INFLATE button — Press.

3-16. AFTER EMERGENCYWATER LANDING

4

If floats are inflated in level flight,there is a possibility that floatswill not align, which will allowright or left forward bag tooscillate. If this occurs, a lowpower descent will align floatbags and stop oscillation.

3. FLOAT ARM switch — Up, guardopen.

WARN

FLIGHT FOLLOWING A WATERLANDING IS PROHIBITED.


Section 4PERFORMA NCE

4-5. HOVER CEILING 4-5-B. OUT-OF-GROUND-EFFECTHOVER

45A. INGROUNDEFFECT Out-of-ground-effect hover performance isHOVER same as basic helicopter.

Subtract 50 pounds (22.68 kilograms) fromIGE hover gross weight for takeoff poweror maximum continuous power.

5/6

B HT-407-FMS-2


SUPPLEMENTHIGH SKID GEAR

407-706-007

CERTIFIED14 FEBRUARY 1996

This supplement shall be attached to Model 407Flight Manual when High Skid Gear kit hasbeen installed.


_________________ Bell Helicopterki *4 ti.11COPYRIGHT NOTICE A Subsidiary 01 Textron Inc.

BELL® HELICOPTER INC.POST OFFICE BOX 482 • FORT WORTH, TEXAS 76101

AND BELL HELICOPTER TEXTRON INC.A DIVISRDN OF TEXTRON CANADA LTD. 14 FEBRUARY 1996

BHT-407-FMS-2

NOTICE PAGE




NP

BHT-407-FMS-2

LOG OF REVISIONS

Original .0 . 14 Feb 96LOG OF PAGES


FLIGHT MANUAL C/D 0i/li 0

Title—NP 0 1—2 0A—B 0

NOTE


A



Original .......... 0 . 14 Feb 96

APPROVED: DATE: /L Fe6 9

1<

CHIEF, FLIGHT TESTFORDIRECTOR — AIRWORTHINESS BRANCHDEPARTMENT OF TRANSPORT

B

BHT-407-FMS-2


Original .0 . 14 Feb 96

C/D


GENERAL INFORMATION

High skid landing gear kit (407-706-007) millimeters) of additional ground clearanceprovides approximately 8.75 inches (222.25 over standard skid gear.

i/il

DOT APPROVED BHT-407FMS-2

[ Section 1LIMITA TIONS

1-6. WEIGHT AND CENTER OF if necessary, to return empty weight CG towithin allowable limits. Refer to Center of

GRAVITY gravity vs weight empty chart in Bl-IT-407-MM-i.

Actual weight change shall be determinedafter kit is installed and ballast readjusted,

[ Section 2NORMAL PROCEDURES

2-10. DESCENT AND WARNINGLANDING

Tail-low run-on landings should be RUN-ON LANDINGS ON OTHERavoided to prevent nosedown pitching. THAN A HARD, FIRM SURFACESHOULD BE EXERCISED WITHCAUTION.


I

No change from basic manual.

1


L Section 4PERFORMA NCE

4-5. HOVER CEILING 4-5-B. OUT- OF-GROUND-EFFECTHOVER

45A. INGROUNDEFFECT Out-of-ground-effect hover performance isHOVER same as basic helicopter.

Subtract 50 pounds (22.68 kilograms) fromIGE hover gross weight for takeoff poweror maximum continuous power.

2

BHT-407-FMS-3


SUPPLEMENT

COPYRIGHT NOTICECOPYRIGHT BELL ® HELICOPTER TEXTRON INC. AND BELL HELICOPTER TEXTRON CANADA LTD. ALL RIGHTS RESERVED

POST OFFICE BOX 482 • FORT WORTH, TEXAS 76101

PARTICLE SEPARATOR206-706-212

CERTIFIED1 MARCH 1996

This supplement shall be attached to Model 407 FlightManual when Particle Separator is installed.

Information contained herein supplements informationof basic Flight Manual. For limitations, Procedures, andPerformance Data not contained in this supplement,consult basic Flight Manual.

REISSUED — 16 DECEMBER 2002

2002

BHT-407-FMS-3

NP 16 DEC 2002




NOTICE PAGE

Manufacturer’s Data portion of this supplement is proprietary to BellHelicopter Textron Inc. Disclosure, reproduction, or use of these data forany purpose other than helicopter operation is forbidden without priorwritten authorization from Bell Helicopter Textron Inc.


BHT-407-FMS-3

NOTE


LOG OF REVISIONS

Original ....................... 0................... 01 MAR 96Revision ...................... 1................... 01 MAY 97Revision ...................... 2................... 04 SEP 98

Revision ......................3 ................... 17 APR 00Reissue........................0 ................... 16 DEC 02

LOG OF PAGES

REVISION REVISIONNO. NO. PAGE PAGE

FLIGHT MANUAL

Title............................................................ 0NP.............................................................. 0A — B ........................................................ 0C/D............................................................. 0i/ii ............................................................... 01 — 38 ....................................................... 039/40 .......................................................... 0

MANUFACTURER’S DATA

41/42 ..........................................................0

16 DEC 2002 A

BHT-407-FMS-3


Original ........................0 ................... 01 MAR 96Revision.......................1 ................... 01 MAY 97Revision.......................2 ................... 04 SEP 98Revision.......................3 ................... 17 APR 00Reissue........................0 ................... 16 DEC 02

B 16 DEC 2002

BHT-407-FMS-3


Original ....................... 0................... 01 MAR 96Revision ...................... 1....................01 MAY 97Revision ...................... 2....................18 MAY 99

Revision ...................... 3 ................... 05 MAY 00Reissue........................ 0 .................... 24 SEP 03

16 DEC 2002 C/D

BHT-407-FMS-3

GENERAL INFORMATION

Bell particle separator kit (206-706-212)consists of particle separator, bleed air tubingand hose, electrical cable and requiredhardware for installation.

This supplement incorporates performanceinformation for various combinations of Bell

kits. It also includes limitations and operatingprocedures made necessary because of kitcombinations. This supplement is notintended to replace approved supplements forother optional equipment, but should be usedin conjunction with such supplements.

16 DEC 2002 i/ii

TC APPROVED BHT-407-FMS-3

16 DEC 2002 1



Particle separator can be removed and theengine air intake screen installed to attainbasic helicopter performance.

1-5. CONFIGURATION

1-5-A. OPTIONAL EQUIPMENT

For operations with part icle separatorinstalled in conjunction with 206-706-208

snow deflector, refer to LIMITATIONS sectionand PERFORMANCE sect ion of snowdeflector supplement (BHT-407-FMS-4).

1-6. WEIGHT AND CENTER OF GRAVITY

Actual weight change shall be determinedafter kit is installed and ballast readjusted, ifnecessary, to return empty weight CG towithin allowable limits.

1

BHT-407-FMS-3 TC APPROVED

2 16 DEC 2002


22-7. BEFORE TAKEOFF

1. PART SEP purge switch – Asrequired.

2-7-A. BEFORE FLIGHT WHEN OPERATING IN SNOW CONDITIONS

1. Thoroughly check cabin roof,transmission cowling, deflectorbaffles and engine air intake areas.All areas checked shall be clean andfree of accumulated snow, slush, andice before each flight.

2. Check engine air plenum chamberthrough plexiglass windows on eachside of inlet cowling for snow, slush,or ice, paying particular attention tofirewalls and rear face of particleseparator. Clean thoroughly beforeeach flight.






3No change from basic manual.


44-2. POWER ASSURANCE

CHECK

Performance is reduced wi th par t ic leseparator installed. This reduction increases

with use of particle separator purge and isprimarily result of bleed air being taken fromengine. A Power assurance check chart(Figure 4-1) is provided to determine if enginecan produce installed power.


16 DEC 2002 3

4

Figure 4-1. Power assurance check

PARTICLE SEPARATOR KITMODEL 407 POWER ASSURANCE CHECK - ROLLS ROYCE 250-C47B ENGINE

HOVER OR LEVEL FLIGHT (85 TO 105 KIAS - NOT TO EXCEED VNE)

48

58

68

78

88

98

108

118

128

138

175 200 225 250 275 300 325 350 375 400 425 450 475 500 525 550 575 600 625 650 675 700 725 750 775

MEASURED GAS TEMPERATURE - DEG.C

35 9085807570656055504540 10095

ENGINE TORQUE - PERCENT

PARTICLE SEPARATOR PURGE OFF GENERATOR LOAD 35 AMPS OR LESS

POWER TURBINE - 100% RPMHEATER / ECS OFF

ANTI-ICE OFF

EXAMPLE: ENTER CHART AT OBSERVED TORQUE (70%)PROCEED VERTICALLY DOWN TO PRESSURE ALTITUDE (6,000 FT.)FOLLOW HORIZONTALLY TO THE RIGHT TO OBSERVED OAT (10 DEG.C)DROP DOWN TO READ MAXIMUM ALLOWABLE MGT (681 DEG.C)

407FMS3 FIG 4-1 REV A.WMF


4 16 DEC 2002

Figure 4-2. Hover ceiling IGE – takeoff power (sheet 1 of 4)


WITH PARTICLE SEPARATOR PURGE ON, REDUCE GROSS WEIGHT BY 55 LBS



PARTICLE SEPARATOR

NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART IN BASIC FLIGHT MANUAL (FIGURE 4-5, 4-5A OR 4-5B IN BHT-407-FM-1)

14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26GROSS WEIGHT - KG x 100

MAXIM

UM OAT

20,000 FT HD

8000

14,00

060

0040

0020

00

12,00

0SEA LE

VEL-20

00

H P - F

T

MAXIM

UM

OAT

10,00

0

16,00

018,00

0

MIN

IMU

M O

AT

EXTE

RN

AL

= 60

00 L

B

MA

X G

WIN

TER

NA

L =

5000

LB

OAT - °C51.7

50

4030

20

10

0-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,00

0

AREA B

M407_FMS-3_FIG 4-2 (1_OF_4).WMF

MAX DEMONSTRATED HD


16 DEC 2002 5





SKID HEIGHT 4 FT (1.2 METER)ANTI-ICE ON

PARTICLE SEPARATOR


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

5

0

-10

-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-3__FIG_4-2_(2_OF_4).WMF

MAX DEMONSTRATED HD


6 16 DEC 2002






PARTICLE SEPARATOR


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

2010

0-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-3__FIG_4-2_(3_OF_4).WMF

MAX DEMONSTRATED HD


16 DEC 2002 7






PARTICLE SEPARATOR


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-3__FIG_4-2_(4_OF_4).WMF

MAX DEMONSTRATED HD


8 16 DEC 2002

Figure 4-3. Hover ceiling IGE – maximum continuous power (sheet 1 of 4)





PARTICLE SEPARATOR


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


MA

XIMU

M O

AT

20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

51.750

4030

20

100

-10-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-3__FIG_4-3_(1_OF_4).EMF

MAX DEMONSTRATED HD


16 DEC 2002 9






PARTICLE SEPARATOR


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-3__FIG_4-3_(2_OF_4).WMF

MAX DEMONSTRATED HD


10 16 DEC 2002






PARTICLE SEPARATOR


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

2010

0-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-3__FIG_4-3_(3_OF_4).EMF

MAX DEMONSTRATED HD


16 DEC 2002 11






PARTICLE SEPARATOR


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-3__FIG_4-3_(4_OF_4).EMF

MAX DEMONSTRATED HD


12 16 DEC 2002

Figure 4-4. Hover ceiling OGE – takeoff power (sheet 1 of 4)



PARTICLE SEPARATOR

NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHARTIN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)



REFER TOFIG. 4-5A

REFER TO FIG. 4-5

M407_FMS-3__FIG_4-4_(1_OF_4).EPS


16 DEC 2002 13






PARTICLE SEPARATOR

NOTE: REFER TO APPROPRIATE HOVER CEILING WIND ACCOUNTABILITY CHART IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)

14,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26

GROSS WEIGHT - KG x 10017,000 FT HD

8000

14,0

00

6000

4000

2000

12,0

00

SEA LEVEL

-200

0

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

-20

50

-10

-30

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5

AREA B

M407_FM S- 3__F IG_4-2_(2_OF_4).WMF


14 16 DEC 2002


HOVER CEILING

OUT OF GROUND EFFECT

TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS

SKID HEIGHT 40 FT (12.2 METER)

HEATER ON

PARTICLE SEPARATOR



REFER TO FIG. 4-5

REFER TO

FIG. 4-5A

M407_FMS-3__FIG_4-4_(3_OF_4).EPS


16 DEC 2002 15


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER AND ANTI-ICE ON

PARTICLE SEPARATOR



HOVER CEILING


REFER TO FIG. 4-5

REFER TO

FIG. 4-5A

M407_FMS-3__FIG_4-4_(4_OF_4).EPS


16 16 DEC 2002

Figure 4-5. Hover ceiling OGE – maximum continuous power (sheet 1 of 4)

MAXIMUM CONTINUOUS POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER AND ANTI-ICE OFF

PARTICLE SEPARATOR

HOVER CEILING



IN BASIC FLIGHT MANUAL (FIGURE 4-5 OR 4-5A IN BHT-407-FM-1)


REFER TO

FIG. 4-5A

REFER TO FIG. 4-5

M407_FMS-3__FIG_4-5_(1_OF_4).EPS


16 DEC 2002 17






PARTICLE SEPARATOR


14,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


8000

14,0

00

6000

4000

2000

12,0

00

SEA LEVEL

-200

0

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

-205

0-10

-30-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5

M407_FMS-3__FIG_4-5_(2_OF_4).WMF


18 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

PARTICLE SEPARATOR

HOVER CEILING





REFER TO

FIG. 4-5A

REFER TO FIG. 4-5

M407_FMS-3__FIG_4-5_(3_OF_4).EPS


16 DEC 2002 19



ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR

HOVER CEILING





REFER TO

FIG. 4-5A

REFER TO FIG. 4-5

M407_FMS-3__FIG_4-5_(4_OF_4).EPS


20 16 DEC 2002

Figure 4-6. Rate of climb takeoff power (sheet 1 of 10)

RATE OF CLIMB


REDUCE RATE OF CLIMB 70 FT/MIN FOR PURGE ON



60 KIASHEATER OFF

PARTICLE SEPARATORPURGE OFF

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10-20

M407_FMS-3__FIG_4-6_(1_OF_10).WMF


16 DEC 2002 21


RATE OF CLIMB





60 KIASHEATER ON


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

20

10

0

-10

-20

-30

M407_FMS-3__FIG_4-6_(2_OF_10).WMF


22 16 DEC 2002


RATE OF CLIMB





60 KIASHEATER OFF


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10 -20

-30-40

40

50

M407_FMS-3__FIG_4-6_(3_OF_10).WMF


16 DEC 2002 23


RATE OF CLIMB





60 KIASHEATER ON


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MAX O

AT LIM

IT

OAT-°C

20

10

0

-10

-20

-30

-40

M407_FMS-3__FIG_4-6_(4_OF_10).WMF


24 16 DEC 2002


RATE OF CLIMB





60 KIASHEATER OFF


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OAT LIM

IT

OAT-°C

30

20

10

0

-10

-20 -30

-40

40

50

M407_FMS-3__FIG_4-6_(5_OF_10).WMF


16 DEC 2002 25


RATE OF CLIMB





60 KIASHEATER ON


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MAX OAT LIM

IT

OAT-°C

20

10

0

-10

-20

-30

-40

M407_FMS-3__FIG_4-6_(6_OF_10).WMF


26 16 DEC 2002


RATE OF CLIMB





60 KIASHEATER OFF


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

-20

-30

-40

40

50

M407_FMS-3__FIG_4-6_(7_OF_10).WMF


16 DEC 2002 27


RATE OF CLIMB





60 KIASHEATER ON


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

20

10

0

-10

-20

-30

-40

M407_FMS-3__FIG_4-6_(8_OF_10).WMF


28 16 DEC 2002


RATE OF CLIMB





60 KIASHEATER OFF


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

-20

-30

-40

40

50

M407_FMS-3__FIG_4-6_(9_OF_10).WMF


16 DEC 2002 29


RATE OF CLIMB





60 KIASHEATER ON


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

20

10

0

-10

-20

-30

-40

M407_FMS-3__FIG_4-6_(10_OF_10).WMF


30 16 DEC 2002

Figure 4-7. Rate of climb maximum continuous power (sheet 1 of 10)407FMS-3-4-7-1.TIF


16 DEC 2002 31

Figure 4-7. Rate of climb maximum continuous power (sheet 2 of 10)

RATE OF CLIMB





60 KIASHEATER ON


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 10

00

ME

TE

RS

MA

X O

AT LIM

ITOAT-°C

20

10

0

-10

-20

-30

-40

M407_FMS-3__FIG_4-7_(2_OF_10).WMF


32 16 DEC 2002


407FMS-3-4-7-3.TIF


16 DEC 2002 33


RATE OF CLIMB





60 KIASHEATER ON


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C20

10

0

-10

-20

-30

-40

M407_FMS-3__FIG_4-7_(4_OF_10).WMF


34 16 DEC 2002


407FMS-3-4-7-5.TIF


16 DEC 2002 35


407FMS-3-4-7-6.TIF


36 16 DEC 2002


407FMS-3-4-7-7.TIF


16 DEC 2002 37


407FMS-3-4-7-8.TIF


38 16 DEC 2002


407FMS-3-4-7-9.TIF



407FMS-3-4-7-10.TIF

16 DEC 2002 39/40

MANUFACTURER’S DATA BHT-407-FMS-3

Section 1SYSTEMS DESCRIPTION

1

This kit is equipped with a PART SEP switchlocated on overhead console. When switch isOFF, engine bleed air is not used to purge

debris from particle separator. When switch isON, engine bleed is used to purge debris.

16 DEC 2002 41/42

BHT-407-FMS-4


SUPPLEMENT

COPYRIGHT NOTICECOPYRIGHT BELL ® HELICOPTER TEXTRON INC.AND BELL HELICOPTER TEXTRON CANADA LTD. ALL RIGHTS RESERVED


SNOW DEFLECTOR206-706-208


This supplement shall be attached to Model 407 FlightManual when Snow Deflector kit is installed.

Information contained herein supplements Informationof basic Flight Manual. For limitations, Procedures, andPerformance Data not contained in this supplement,consult basic Flight Manual


2002

BHT-407-FMS-4

NP 16 DEC 2002




NOTICE PAGE

Manufacturer’s Data portion of this supplement is proprietary to BellHelicopter Textron Inc. Disclosure, reproduction, or use of these data forany purpose other than helicopter operation is forbidden without priorwritten authorization from Bell Helicopter Textron Inc.


BHT-407-FMS-4

NOTE


LOG OF REVISIONS

Original ....................... 0....................01 MAR 96Revision ...................... 1....................27 MAY 97Revision ...................... 2....................04 SEP 98

Revision ......................3 ................... 17 APR 00Reissue........................0 ................... 16 DEC 02

LOG OF PAGES


FLIGHT MANUAL

Title............................................................ 0NP.............................................................. 0A — B ........................................................ 0

C/D .............................................................0i/ii ...............................................................01 — 76........................................................0

16 DEC 2002 A



Original ........................0 ...................01 MAR 96Revision.......................1 ...................27 MAY 97Revision.......................2 ...................04 SEP 98

Revision ...................... 3 ................... 17 APR 00Reissue ....................... 0.................... 16 DEC 02

B 16 DEC 2002

BHT-407-FMS-4


Original ....................... 0................... 01 MAR 96Revision ...................... 1................... 27 MAY 97Revision ...................... 2................... 18 MAY 99

Revision ......................3 ................... 05 MAY 00Reissue........................0 ....................24 SEP 03

16 DEC 2002 C/D

BHT-407-FMS-4

GENERAL INFORMATION

Snow deflector kit (206-706-208) consists oftwo deflectors that mount on either side of

transmission fairing, just forward of engineair inlets.

16 DEC 2002 i/ii

16 DEC 2002 1




Snow deflector kit shall be installed foroperation in falling or blowing snow. Theymay be installed with basic inlet screen orparticle separator kit (206-706-212).


Actual weight changes shall be determinedafter kit is installed and ballast readjusted, ifnecessary, to return empty weight CG to withinallowable limits. Refer to Center of gravity vsweight empty chart in BHT-407-MM-1.

1-11. AMBIENT TEMPERATURES

Snow deflectors shal l be removed foroperations above 30°C (86°F).

1-22. SNOW OPERATION

For operation in falling or blowing snow, thefollowing limits apply:

Hover flight in falling and/or blowing snow is limited to 15 minute duration after which helicopter shall be landed and checked for snow and/or ice accumulation.

Flight operations are prohibited when visibility in falling or blowing snow is less than one-half (1/2) statute mile.


2 16 DEC 2002



2-3-A. EXTERIOR CHECK

2-3-A-1. OPERATION IN FALLING OR BLOWING SNOW

1. Thoroughly check cabin roof, transmission fairing, deflector baffles, and engine air intake areas. All areas checked shall be clean and free of accumulated snow, slush, and ice before each flight.

NOTE

Due to reduced performance at highertemperatures, it is recommended thatsnow deflectors be removed above20°C (68°F).

2. Particle separator kit (if installed), check engine air plenum chamber through plexiglass windows on each side of inlet cowling for snow, slush, or ice, paying particular attention to firewalls and rear face of particle separator. Clean thoroughly before each flight.

2-3-A-2. AFTER EXITING HELICOPTER

WARNING

FAILURE TO INSTALL ENGINEINTAKE COVERS COULD ALLOWFALLING/BLOWING SNOW TO ENTERTHE PARTICLE SEPARATOR PLENUM(IF INSTALLED).

Install protective covers (engine intake,exhaust, and pitot tube).




16 DEC 2002 3


44-1. INTRODUCTION

Refer to appropriate performance charts inaccordance with optional equipment installed.

NOTEDue to reduced performance at highertemperatures, it is recommended thatsnow deflectors be removed above20°C (68°F).

4-2. POWER ASSURANCE CHECK

Power assurance check chart Figure 4-1.

Performance is reduced with snow bafflesinstalled. Power assurance check chart(Figure 4-1) is provided to determine if enginecan produce insta l led power. Powerassurance check shall be conducted in levelflight only. This chart is valid for both basicinlet with snow deflectors installed and snowdeflectors with particle separator installed.PARTICLE SEP PRG switch (if installed) shallbe OFF when performing a power assurancecheck.

4-5. HOVER CEILING

4-5-A. HOVER CEILING IN-GROUND-EFFECT

Hover ceiling IGE (takeoff power) charts arepresented in Figure 4-2, and Hover ceiling IGE(maximum continuous power) charts arepresented in Figure 4-3.

4-5-B. HOVER CEILING OUT-OF-GROUND-EFFECT

Hover ceiling OGE (takeoff power) charts arepresented in Figure 4-4, and Hover ceilingOGE (maximum continuous power) charts arepresented in Figure 4-5.


4-7-A. RATE OF CLIMB

Rate of climb (takeoff power) charts arepresented in Figure 4-6, Rate of cl imb(maximum continuous power) charts arepresented in Figure 4-7, Rate of climb (takeoffpower) (part ic le separator) charts arepresented in Figure 4-8, and Rate of climb(maximum continuous power) (particleseparator) charts are presented in Figure 4-9.


4 16 DEC 2002

Figure 4-1. Power assurance check chart

SNOW DEFLECTOR KITMODEL 407 POWER ASSURANCE CHECK - ROLLS ROYCE 250-C47B ENGINE

LEVEL FLIGHT (85 TO 105 KIAS - NOT TO EXCEED VNE)

48

58

68

78

88

98

108

118

128

175 200 225 250 275 300 325 350 375 400 425 450 475 500 525 550 575 600 625 650 675 700 725 750 775

MEASURED GAS TEMPERATURE - DEG.C

35 9085807570656055504540 10095

ENGINE TORQUE - PERCENT

PARTICLE SEPARATOR PURGE OFFGENERATOR LOAD 35 AMPS OR LESS

POWER TURBINE - 100% RPMHEATER / ECS OFF

ANTI-ICE OFF

EXAMPLE: ENTER CHART AT OBSERVED TORQUE (70%)PROCEED VERTICALLY DOWN TO PRESSURE ALTITUDE (6,000 FT.)FOLLOW HORIZONTALLY TO THE RIGHT TO OBSERVED OAT (10 DEG.C)DROP DOWN TO READ MAXIMUM ALLOWABLE MGT (721 DEG.C)

407FMS4 FIG 4-1.WMF

16 DEC 2002 5






BASIC INLETSNOW DEFLECTOR


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


MA

XIM

UM

OA

T20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

3020

10

0

-10

-30-20

-40 REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-2_(1_OF_8).WMF

MAX DEMONSTRATED HD


6 16 DEC 2002





SNOW DEFLECTOR


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10

-30

-20-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-2_(2_OF_8).WMF

MAX DEMONSTRATED HD

16 DEC 2002 7








14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

2010

0-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-2_(3_OF_8).WMF

MAX DEMONSTRATED HD


8 16 DEC 2002







14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-2_(4_OF_8).WMF

MAX DEMONSTRATED HD

16 DEC 2002 9







PARTICLE SEPARATORSNOW DEFLECTOR


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


MA

XIM

UM

OA

T20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

3020

100

-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-2_(5_OF_8).WMF

MAX DEMONSTRATED HD


10 16 DEC 2002








14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10

-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-2_(6_OF_8).WMF

MAX DEMONSTRATED HD

16 DEC 2002 11









14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

2010

0-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-2_(7_OF_8).WMF

MAX DEMONSTRATED HD


12 16 DEC 2002








14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-2_(8_OF_8).WMF

MAX DEMONSTRATED HD

16 DEC 2002 13








14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

30 20

10

0-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-3_(1_OF_8).EMF

MAX DEMONSTRATED HD


14 16 DEC 2002





SNOW DEFLECTOR


14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-3_(2_OF_8).WMF

MAX DEMONSTRATED HD

16 DEC 2002 15








14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

2010

0-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-3_(3_OF_8).EMF

MAX DEMONSTRATED HD


16 16 DEC 2002







14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10-30

-20-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-3_(4_OF_8).EMF

MAX DEMONSTRATED HD

16 DEC 2002 17









14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

30 20

100

-10-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-3_(5_OF_8).EMF

MAX DEMONSTRATED HD


18 16 DEC 2002








14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10-30

-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-3_(6_OF_8).WMF

MAX DEMONSTRATED HD

16 DEC 2002 19









14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

2010

0-10

-30-20

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-3_(7_OF_8).EMF

MAX DEMONSTRATED HD


20 16 DEC 2002








14,000 FT HD

17,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


20,000 FT HD

8000

14,0

0060

0040

0020

00

12,0

00SEA L

EVEL-2

000

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

0018,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

50

-10-30

-20-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5B

REFER TO FIG. 4-5

20,0

00

AREA B

M407_FMS-4__FIG_4-3_(8_OF_8).EMF

MAX DEMONSTRATED HD

16 DEC 2002 21








14,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


8000

14,0

00

6000

4000

2000

12,0

00

SEA LEVEL

-200

0

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

-2010

0

-10

-30

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5

AREA B

MA

X O

AT

3020

M407_FMS-4__FIG_4-4_(1_OF_8).WMF


22 16 DEC 2002





SNOW DEFLECTOR


14,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


8000

14,0

00

6000

4000

2000

12,0

00

SEA LEVEL

-200

0

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

-20

50

-10 -30

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5

AREA B

M407_FMS-4__FIG_4-4_(2_OF_8).WMF

16 DEC 2002 23



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

BASIC INLET

SNOW DEFLECTOR

HOVER CEILING


REFER TO

FIG. 4-5A

REFER TO FIG. 4-5



M407_FMS-4__FIG_4-4_(3_OF_8).EPS


24 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET

SNOW DEFLECTOR

HOVER CEILING


REFER TO

FIG. 4-5A

REFER TO FIG. 4-5



M407_FMS-4__FIG_4-4_(4_OF_8).EPS

16 DEC 2002 25









14,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


8000

14,0

00

6000

4000

2000

12,0

00

SEA LEVEL

-200

0

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

-20

10

0

-10

-30

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5

AREA B

MA

X O

AT

3020

M407_FMS-4__FIG_4-4_(5_OF_8).WMF


26 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

PARTICLE SEPARATOR

SNOW DEFLECTOR

HOVER CEILING





REFER TO

FIG. 4-5A

REFER TO FIG. 4-5

M407_FMS-4__FIG_4-4_(SHEET_6).EPS

16 DEC 2002 27



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

PARTICLE SEPARATOR

SNOW DEFLECTOR

HOVER CEILING





REFER TO

FIG. 4-5A

REFER TO FIG. 4-5



28 16 DEC 2002








14,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


8000

14,0

00

6000

4000

2000

12,0

00

SEA LEVEL

-200

0

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

-205

0-10

-30

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5

M407_FMS-4__FIG_4-4_(8_OF_8).WMF

16 DEC 2002 29




ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET

SNOW DEFLECTOR

HOVER CEILING




REFER TO

FIG. 4-5A

REFER TO FIG. 4-5



30 16 DEC 2002





SNOW DEFLECTOR


14,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


8000

14,0

00

6000

4000

2000

12,0

00

SEA LEVEL

-200

0

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

-205

0-10

-30

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5

M407_FMS-4__FIG_4-5_(2_OF_8).WMF

16 DEC 2002 31




ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

BASIC INLET

SNOW DEFLECTOR

HOVER CEILING




REFER TO

FIG. 4-5A

REFER TO FIG. 4-5



32 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET

SNOW DEFLECTOR

HOVER CEILING




REFER TO

FIG. 4-5A

REFER TO FIG. 4-5


16 DEC 2002 33




ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR

SNOW DEFLECTOR

HOVER CEILING





REFER TO

FIG. 4-5A

REFER TO FIG. 4-5



34 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

PARTICLE SEPARATOR

SNOW DEFLECTOR

HOVER CEILING





REFER TO

FIG. 4-5A

REFER TO FIG. 4-5


16 DEC 2002 35




ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

PARTICLE SEPARATOR

SNOW DEFLECTOR

HOVER CEILING





REFER TO

FIG. 4-5A

REFER TO FIG. 4-5



36 16 DEC 2002








14,000 FT HD

-40 -20 0 20 40 60

OAT - °C32 36 40 44 48 52 56 60


16 18 20 22 24 26


8000

14,0

00

6000

4000

2000

12,0

00

SEA LEVEL

-200

0

H P -

FT

MA

XIM

UM

OA

T

10,0

00

16,0

00

MIN

IMU

M O

AT

EX

TE

RN

AL

= 6

000

LB

MA

X G

WIN

TE

RN

AL

= 5

000

LB

OAT - °C

-205

0-10

-30

-40

REFER TO FIG. 4-5A

REFER TO FIG. 4-5

M407_FMS-4__FIG_4-5_(8_OF_8).WMF

16 DEC 2002 37


Figure 4-6. Rate of climb (takeoff power) (sheet 1 of 10)

407FMS-4-4-6-1


38 16 DEC 2002


407FMS-4-4-6-2

16 DEC 2002 39



407FMS-4-4-6-3


40 16 DEC 2002


407FMS-4-4-6-4

16 DEC 2002 41



407FMS-4-4-6-5


42 16 DEC 2002


407FMS-4-4-6-6

16 DEC 2002 43



407FMS-4-4-6-7


44 16 DEC 2002


407FMS-4-4-6-8

16 DEC 2002 45



RATE OF CLIMB





SNOW DEFLECTOR

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MAX OAT

LIMITOAT-°C

30

20

10

0

-10

-20-30

-40

M407_FMS-4__FIG_4-6_(9_OF_10).WMF


46 16 DEC 2002


407FMS-4-4-6-10

16 DEC 2002 47


Figure 4-7. Rate of climb (maximum continuous power) (sheet 1 of 10)

407FMS-4-4-7-1


48 16 DEC 2002


407FMS-4-4-7-2

16 DEC 2002 49



407FMS-4-4-7-3


50 16 DEC 2002


407FMS-4-4-7-4

16 DEC 2002 51



407FMS-4-4-7-5


52 16 DEC 2002


407FMS-4-4-7-6

16 DEC 2002 53



RATE OF CLIMB





SNOW DEFLECTOR

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

30

20

10

0

-10

-20

-30

-40

MA

X OA

T

LIMIT

0

-20

-40

M407_FMS-4__FIG_4-7_(7_OF_10).WMF


54 16 DEC 2002


407FMS-4-4-7-8

16 DEC 2002 55



407FMS-4-4-7-9


56 16 DEC 2002


407FMS-4-4-7-10

16 DEC 2002 57


Figure 4-8. Rate of climb (takeoff power) (particle separator) (sheet 1 of 10)407FMS-4-4-8-1


58 16 DEC 2002

Figure 4-8. Rate of climb (takeoff power) (particle separator) (sheet 2 of 10)

407FMS-4-4-8-2

16 DEC 2002 59



407FMS-4-4-8-3


60 16 DEC 2002


407FMS-4-4-8-4

16 DEC 2002 61



407FMS-4-4-8-5


62 16 DEC 2002


407FMS-4-4-8-6

16 DEC 2002 63



407FMS-4-4-8-7


64 16 DEC 2002


407FMS-4-4-8-8

16 DEC 2002 65



RATE OF CLIMB





60 KIASHEATER OFF

PARTICLE SEPARATOR - PURGE OFFSNOW DEFLECTOR

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MAX OAT

LIMIT

OAT-°C

30

20

10

0

-10

-20

-30

-40

M407_FMS-4__FIG_4-8_(9_OF_10).WMF


66 16 DEC 2002


407FMS-4-4-8-10

16 DEC 2002 67


Figure 4-9. Rate of climb (maximum continuous power) (particle separator) (sheet 1 of 10)

407FMS-4-4-9-1


68 16 DEC 2002


407FMS-4-4-9-2

16 DEC 2002 69



407FMS-4-4-9-3


70 16 DEC 2002


407FMS-4-4-9-4

16 DEC 2002 71



407FMS-4-4-9-5


72 16 DEC 2002


407FMS-4-4-9-6

16 DEC 2002 73



407FMS-4-4-9-7


74 16 DEC 2002


407FMS-4-4-9-8

16 DEC 2002 75



407FMS-4-4-9-9


76 16 DEC 2002


407FMS-4-4-9-10


SUPPLEMENT

CARGO HOOK206-706-341

AND407-704-023


This supplement shall be attached to the BHT-407-FM-1when the Cargo Hook kit has been installed.

Information contained herein supplements information inthe basic Flight Manual. For Limitations, Procedures, andPerformance Data not contained in this supplement, refer tothe basic Flight Manual.

BHT-407-FMS-5

REISSUE — 15 NOVEMBER 2007REVISION 1 — 25 MARCH 2008


ALL RIGHTS RESERVED

NOTICE PAGE







NP———15 NOV 2007

BHT-407-FMS-5

These data are proprietary to Bell Helicopter Textron Inc. Disclosure,reproduction, or use of these data for any purpose other than helicopteroperation or maintenance is forbidden without prior written authorizationfrom Bell Helicopter Textron Inc.


Original ..................... 0...................... 14 FEB 96Revision .................... 1...................... 04 SEP 98

Reissue......................0 ..................... 15 NOV 07Revision.....................1 .....................25 MAR 08

25 MAR 2008—Rev. 1———A/B

BHT-407-FMS-5

NOTE


LOG OF REVISIONS

LOG OF PAGES


FLIGHT MANUAL

Title.............................................................1NP ...............................................................0A/B..............................................................1C/D..............................................................1E/F ..............................................................1i/ii ................................................................11 ..................................................................02 ..................................................................13 – 10 ..........................................................0

APPROVED DATE


25 MAR 2008—Rev. 1———C/D

BHT-407-FMS-5

Original ..................... 0...................... 14 FEB 96Revision .................... 1...................... 04 SEP 98

Reissue......................0 ..................... 15 NOV 07Revision ....................1 .....................25 MAR 08



Original ..................... 0...................... 14 FEB 96Revision .................... 1......................18 MAY 99

Reissue......................0 ..................... 21 DEC 07Revision ....................1 ..................... 05 MAY 08

25 MAR 2008—Rev. 1———E/F

BHT-407-FMS-5

BHT-407-FMS-5

GENERAL INFORMATION

25 MAR 2008—Rev. 1———i/ii

Installation of Cargo Hook kit (206-706-341) or Cargo Hook Retrofit kit (407-704-023) with anapproved hook assembly adds capability of transporting external cargo. Kit contains electricaland manual releases, both operated from the pilot seat. Cargo hook is located at FS 121.0 (3073mm).

Approved assemblies may have a spring-loaded keeper (P/N 17149-9 or 528-010-ALL) or akeeperless hook (P/N 528-023-ALL) design.

Cargo hook kit will permit operator to use helicopter for transportation of external cargo.


15 NOV 2007———1



Operation of helicopter with no load onexternal cargo suspension hook is authorizedunder standard airworthiness certificatewithout removing unit from helicopter.

With a load attached to suspension assembly,operations shall be conducted in accordancewith appropriate operating rules for externalloads.


Actual weight change shall de determinedafter cargo hook is installed and ballastreadjusted, if necessary, to return emptyweight CG to within allowable limits. Refer toCenter of Gravity vs Weight Empty Chart inthe BHT-407-MM-1.

CAUTION

LOADS THAT RESULT IN GROSSWEIGHTS ABOVE 5000 POUNDS(2268 KG) SHALL BE CARRIED ONCARGO HOOK AND SHALL BEJETTISONABLE.

Maximum gross weight of helicopter andexternal load operations is 6000 pounds(2724 kg).

Maximum cargo hook load is 2650 pounds(1202 kg).

Refer to the BHT-407-FM-1 for Gross WeightCenter of Gravity Limits charts for externalcargo operations.

1-7. AIRSPEED

VNE with external cargo load is 100 KIAS.

CAUTION

AIRSPEED WITH EXTERNAL CARGOIS LIMITED BY CONTROLLABILITY.CAUTION SHOULD BE EXERCISEDWHEN CARRYING EXTERNALCARGO, AS HANDLINGCHARACTERIST ICS MAY BEAFFECTED BY SIZE, WEIGHT, ANDSHAPE OF CARGO LOAD.

Light weight, high drag loads require a swivelconnector between cargo hook and sling toprevent unstable oscillations in flight above20 KIAS.


Refer to Figure 1-3 for decals.


2———Rev. 1—25 MAR 2008

Figure 1-3. Placards and Decals

407-FMS-5_0001

Kit 206-706-341-109

Location: On cargo hook roller beam.

Kits 206-706-341-141 and 407-704-023-103

Location: On cargo hook beam.


15 NOV 2007———3



2-2-A. GROUND CREW INSTRUCTIONS

Instruct ground crew member to dischargehelicopter static electricity before attachingcargo by touching airframe with ground wire,or, if metal sling is used, hookup ring can bestruck against cargo hook. If contact has beenlost after initial grounding, helicopter shall beelectrically regrounded and, if possible,contact mainta ined unt i l hookup iscompleted.

1. Cargo hook — Condition and security.Instruct ground personnel to checkprimary load ring and secondary loadring for condition and proper size(Table 2-1). Check for correct rigging(Figure 2-1 and Figure 2-2).

WARNING

USE OF INAPPROPRIATELY SIZEDLOAD RINGS MAY RESULT IN LOADHANG-UP WHEN LOAD RING IS TOOSMALL OR INADVERTENT LOAD

RELEASE IF LOAD RING IS TOOLARGE.

CAUTION

THE EXAMPLES SHOWN INFigure 2-1 THROUGH Figure 2-3 ARENOT INTENDED TO REPRESENTALL POSSIBILITIES. IT IS THERESPONSIBIL ITY OF THEOPERATOR TO MAKE SURE THEHOOK WILL FUNCTION PROPERLYWITH THE R IGGING. SOMECOMBINATIONS OF SMALLPRIMARY RINGS AND LARGESECONDARY RINGS COULD CAUSEFOULING DURING RELEASE.

2. Check that only one primary ring iscaptured in the load beam and onlyone secondary ring with correctcross-section dimension is capturedin the primary ring. Additional rings,slings, or shackles shall be attachedto the secondary load ring (Figure 2-1through Figure 2-3).

Table 2-1: Cargo Hook Ring Sizes

Primary RingInside Diameter

Primary RingCross-section

Secondary RingMaximum Cross-section

CARGO HOOK P/N 17149-6

1.50 to 1.68 inches(38.10 to 42.67 mm)

0.75 inch(19.05 mm)

0.438 inch(11.12 mm)

CARGO HOOK P/N 528-010-ALL

1.50 to 1.87 inches(38.10 to 47.50 mm)

0.875 inch(22.22 mm)

0.625 inch(15.88 mm)

CARGO HOOK P/N 528-023-ALL (KEEPERLESS HOOK)

Refer to Figure 2-4.


4———15 NOV 2007


2-3-B. EXTERIOR CHECK

Cargo suspension assembly — Condition andsecurity.

Cargo sling — Condition, proper length.

2-4. INTERIOR AND PRESTARTCHECK

2-4-A. INTERIOR CHECK(SPRING-LOADED KEEPER)

1. CARGO HOOK circuit breaker — In.

2. Cyclic CARGO RELEASE switch —Press and release; pull down on cargohook; hook should open. Releasecargo hook; hook should close andlock.

3. EMERG CARGO RELEASE PULLhandle — Pull and hold; pull down oncargo hook; hook should open. Pushhandle in; hook should close and lock.

2-4-B. INTERIOR CHECK (KEEPERLESSHOOK)

1. Cargo hook circuit breaker — In.

2. Cyclic CARGO RELEASE switch —Press and release; hook should open.Push hook up manually and verify thatit latches closed.

3. EMERG CARGO RELEASE PULLhandle — Pull and hold; pull down oncargo hook; hook should open. Pushhandle in; push hook up manually andverify that it latches closed.

2-7. BEFORE TAKEOFF

CARGO HOOK circuit breaker — In.

EMERG CARGO RELEASE PULL handle — In.

2-8. TAKEOFF

1. Hover helicopter at sufficient height toallow ground crew member todischarge static electricity and attachcargo sling to cargo hook.

NOTE

For keeper less hook, pr io r toexternal load attachment, ensurehook is in OPEN posi t ion bymomentarily depressing cycl icCARGO RELEASE switch. Groundpersonnel must ensure that loadbeam is latched closed after load isapplied.

2. Ascend vertically, directly over load,then slowly lift load from surface.

3. Pedals — Check for adequatedirectional control.

4. Hover power — Check TORQUErequired to hover with external load.

5. Take off into wind, if possible,allowing adequate sling loadclearance over obstacles.


NOTE

Control movements should be madesmoothly and kept to a minimum toprevent oscillation of sling load.

EMERG CARGO RELEASE PULLhandle will function regardless ofCARGO RELEASE switch position.

1. AIRSPEED — Within limits foradequate controllability of helicopterload combination.

2. Flight path — As planned to avoidflight with external load over anyperson, vehicle, or structure.


15 NOV 2007———5


1. Flight path and approach angle — Asrequired for wind direction andobstacle clearance.

2. Execute approach to a hover with loadclear of surface. When stabilized at ahover, descend slowly until loadcontacts surface. Maintain tension onsling.

3. Cyclic CARGO RELEASE switch —Press to release sling from hook.


6———15 NOV 2007

Figure 2-1. External Load Rigging (Cargo Hook P/N 17149-6)


15 NOV 2007———7

Figure 2-2. External Load Rigging (Cargo Hook P/N 528-010-ALL)

CORRECT RIGGING

INCORRECT RIGGING INCORRECT RIGGING

LOAD

PRIMARYRING I.D.

MAXIMUM CROSS SECTION

OF PRIMARY RING

MAXIMUM CROSS SECTION

OF SECONDARY RING

MULTIPLE RINGS

ON LOAD BEAM

MULTIPLE RINGS

ON LOAD BEAM

407_FMS_5_0002

1 Refer to Table 2-1.

NOTE

1

1

1


8———15 NOV 2007

Figure 2-3. External Load Rigging (Cargo Hook P/N 528-023-ALL)

407_FMS_5_0004

PRIMARY RING

LOAD

SECONDARY RING

1 Refer to Figure 2-4.

NOTE

1


15 NOV 2007———9

Figure 2-4. Cargo Hook P/N 528-023-ALL Ring Sizes

2.16 IN. (55 mm)

MIN.

SECONDARY RING

407_FMS_5_0003

PRIMARY RING

1.97 IN. (50 mm)

MIN. 1.02 IN. (26 mm)

MAX.


10———15 NOV 2007


33-11. CARGO FAILS TO RELEASE

ELECTRICALLY

WARNING

EMERG CARGO RELEASE PULLHANDLE WILL FUNCTION

REGARDLESS OF CARGO RELEASESWITCH POSITION.

In the event the cargo hook will not releasethe sling when the cyclic CARGO RELEASEswitch is pressed, proceed as follows:

1. Maintain tension on sling.

2. Pull EMERG CARGO RELEASE PULLhandle to release load.


44-5. HOVER CEILING

Refer to the BHT-407-FM-1 for out of groundeffect hover performance.

There is no change from the BHT-407-FM-1performance with no load attached to thecargo hook.

Performance may be affected by size andshape of external load.


55-2. EMPTY WEIGHT CENTER OF

GRAVITY

Load on hook is at FS 121.0 (3073 mm).

B HT-407-FMS-6

407ROTORCRAFT

FLIGHT MANUAL

SUPPLEMENTAUXILIARY FUEL KIT

407-706-011CERTIFiED

20 MARCH 1996

This supplement shall be attached to Model 407Flight Manual when AUXILIARY FUEL KIT kithas been installed.


_______________ Bell Helicopterk1*1i.1TlCOPYRIGHT NOTICE A Subsidiary 01 Textron inc.

COPYRIGHT 1996BELL HELICOPTER INC. POST OFFICE BOX 482 • FORT WORTH, TEXAS 18101

AND BELL HELICOPTER TEXTRON INC.A DIVION OF TEXTRON CANADA LTD. 20 MARCH 1 996

BHT-407-FMS-6

NOTICE PAGE




NP

BHT-407-FMS-6

LOG OF REVISIONS

Original .0 . 20 MAR 96LOG OF PAGES

REViSION REVISIONPAGE NO. PAGE NO.

FLIGHT MANUAL C/D 0i/il 0

Title—NP 0 1—6 0A—B 0

NOTE

Revised text is indicated by a black vertical line, insert latest revision pages; dispose ofsuperseded pages.

A


LOG OF APPROVED REViSIONS

Original .0 . 20 MAR 96

APPROVED: DATE: Z0 p1&& 6

-CHIEF, FLIGHT TESTFORDIRECTOR — AIRWORTHINESS BRANCHDEPARTMENT OF TRANSPORT

B

BHT-407-FMS-6


Original .0 . 20 MAR 96

C/D


GENERAL INFORMATION

Auxiliary fuel kit (407-706-011) consists of fuel tank, tubing, electrical wiring, micro-switchand hardware for installation. Removable 19 U.S. gallons (71.9 liters) fuel tank is mountedin baggage compartment to aft bulkhead. Fuel transfers between main aft fuel cell andauxiliary fuel cell for filling and emptying by gravity.

i/u


Section 1LIMITA TIONS


GRAVITY gravity vs weight empty chart in BHT-407-MM-i.

Actual weight changes shall be determinedafter kit is installed and ballast readjusted,

L Section 2NORMAL PROCEDURES

2-2. FLIGHT PLANNING 2-3. PREFLIGHT CHECK

Baggage compartment — Check auxiliaryWith auxiliary fuel tank installed, full fuel fuel tank security and condition.indication is approximately 1005 lbs. (146.9U.S. gallons/556 liters) of Jet A.



Section 4

2


PERFORMA NCE


Section 5

BHT-407-FMS-6

WEIGHT AND BALANCE

5-7. FUEL LOADING 2. Critical fuel for computing most aftuseful load is 146.9 U.S. gallons (556liters).

Longitudinal center of gravity of fuel shifts Fuel loading tables (figure 5-2) list usableas it is consumed (Figure 5-1). Extreme fuel quantities, weight, and moments ineffects of fuel consumption on helicopter both U.S. and metric units.center of gravity for standard fuel systemare as follows:

1. Critical fuel for computing mostforward useful load is 74.5 U.S.gallons (282.0 liters).

3

BHT-407-FMS-6

4

Figure 5-1. Fuel Center of Gravity - Auxiliary Fuel.

F- —

U)

0Cu( 90

80

70U-

60

50

40

30

20

10

0114

160———--————--———————--——--———-——— — — — — — — — — — — — — — —150 — — — — — — — — —

:120

11111 11111; 11111110

- -——- /-s——-—- -—-.--———- -———- -———-

116 118 120 122 124 126 128 130 132 134 136 136

Fuselage Station-Inches

600

550

500

450

400

U) 350a)

300a,

CL. 250

200

150

100

50

02900

VV—--------z-——------—

— — — — — —— — — — —

-———

-—--———--— —— —

- —.---_- — - - - —

-3000 3100 3200 3300

Fuselage Station-Millimeters3400 3500

407-FS-5-t

AUXILIARY FUEL LOADING (U.S.)

* MOST FORWARD FUEL C.G.D CRITICAL FUEL FOR MOST FORWARD C.G. CONDITIONA FULL FUEL— CRITICAL FUEL FOR MOST AFT C.G. CONDITION

BHT-407-FMS-6

Figure 5-2. Auxiliary Fuel Loading - Sheet 1 of 2

5

FUEL LOADING TABLE (U.S.)JP-4 LONGITUDINAL

QUANTITY WE1GHT C.G. MOMENT

(U.S. GAL) (LBS) (IN) (IN-LBS)

JP-5, JP-8 LONGITUDINALQUANTITY WEIGHT C.G. MOMENT

(U.S. GAL) (LBS) (IN) (IN-LBS)5 32.5 133.7 4345

10 65.0 135.0 8,77515 97.5 135.9 13,25020 130.0 136.4 17,73225 162.5 136.7 22,214

28.4 184.6 137.0 25,290

30 195.0 134.3 26,18935 227.5 127.8 29,07540 260.0 122.9 31,95445 292.5 119.1 34,837

50 325.0 116.0 37,70* 50.6 328.9 115.7 38,054

55 357.5 116.1 41,50660 390.0 116.2 45,31865 422.5 116.2 49,09570 455.0 116.1 52,826

074.5 484.3 116.1 56,22775 487.5 116.3 56,69680 520.0 118.0 61,36085 552.5 119.6 66,07990 585.0 121.0 70,78595 617.5 122.3 75,520

100 650.0 123.4 80,210105 682.5 124.5 84,971110 715.0 125.5 89,733

115 747.5 126.5 94,559120 780.0 127.5 99,450

125 812.5 128.5 104,406130 845.0 129.4 109,343135 877.5 130.2 114,251140 910.0 131.0 119,210145 942.5 131.7 124,127

A 146.9 954.9 132.0 126,047

5 34.0 133.7 454610 68.0 135.0 9,18015 102.0 135.9 13,86220 136.0 136.4 18,55025 170.0 136.7 23,239

28.4 193.1 137.0 26,45530 204.0 134.3 27,39735 238.0 127.8 30,41640 272.0 122.9 33,42945 306.0 119.1 36,44550 340.0 116.0 39,440

* 50.6 344.1 115.7 39,81255 374.0 116.1 43,42160 408.0 116.2 47,41065 442.0 116.2 51,36070 476.0 116.1 55,264

074.5 506.6 116.1 58,81675 510.0 116.3 59,31380 544.0 118.0 64,19285 578.0 119.6 69,12990 612.0 121.0 74,05295 646.0 122.3 79,006

100 680.0 123.4 83,912105 714.0 124.5 88,893110 748.0 125.5 93,874115 782.0 126.5 98,923120 816.0 127.5 104,040

125 850.0 128.5 109,225130 884.0 129.4 114,390135 918.0 130.2 119,524140 952.0 131.0 124,712145 986.0 131.7 129,856

A 146.9 998.9 132.0 131,855

BHT-407-FMS-6

AUXILIARY FUEL LOADING (METRIC)

6

* MOST FORWARD FUEL C.G.U CRITICAL FUEL FOR MOST FORWARD C.G. CONDITIONA FULL FUEL— CRITICAL FUEL FOR MOST AFT C.G. CONDITION

Figure 5-2. Auxiliary Fuel Loading - Sheet 2 of 2

FUEL LOADING TABLE (METRIC)

JP-4 LONGITUDINAL

QUANTITY WEIGHT C.G. MOMENT

(LITERS) (kg) (mm) (kg-mm/100)

JP-5, JP-8 LONGITUDINALQUANTITY WEIGHT C.G. MOMENT

(LITERS) (kg) (mm) (kg-mm/100)15 11.7 3389 397

30 23.4 3415 79945 35.0 3439 120460 46.7 3455 161375 58.4 3465 202490 70.1 3472 2434

105 81.8 3478 2845107.5 83.7 3479 2912

120 93.5 3352 3134135 105.1 3228 3393150 116.8 3129 3655165 128.5 3049 3918180 140.2 2982 4181

*191.6 149.2 2938 4383195 151.9 2940 4466210 163.6 2949 4825225 175.2 2951 5170240 186.9 2953 5519255 198.6 2950 5859270 210.3 2948 6200

D282.0 219.7 2949 6479285 222.0 2956 6562300 233.7 2991 6990315 245.3 3024 7418330 257.0 3054 7849345 268.7 3082 8281360 280.4 3107 8712375 292.1 3130 9143390 303.8 3152 9576405 315.4 3172 10004420 327.1 3192 10441435 338.8 3213 10886450 350.5 3234 11335465 362.2 3253 11782480 373.9 3272 12234495 385.5 3290 12683510 397.2 3306 13131525 408.9 3322 13584540 420.6 3337 14035555 432.3 3351 14486

A556.1 433.1 3352 14518

15 12.2 3389 41330 24.4 3415 83345 36.7 3439 126260 48.9 3455 168975 61.1 3465 211790 73.3 3472 2545

105 85.6 3478 2977107.5 87.6 3479 3048

120 97.8 3352 3278135 110.0 3228 3551150 122.2 3129 3824165 134.4 3049 4098180 146.7 2982 4375

* 191.6 156.1 2938 4586195 158.9 2940 4672210 171.1 2949 5046225 183.3 2951 5409240 195.6 2953 5776255 207.8 2950 6130270 220.0 2948 6486

UIJ 282.0 229.8 2949 6777285 232.2 2956 6864300 244.5 2991 7313315 256.7 3024 7763330 268.9 3054 8212345 281.1 3082 8664360 293.3 3107 9113375 305.6 3130 9565390 317.8 3152 10017405 330.0 3172 10468420 342.2 3192 10923435 354.5 3213 11390450 366.7 3234 11859465 378.9 3253 12326480 391.1 3272 12797495 403.3 3290 13269510 415.6 3306 13740525 427.8 3322 14212540 440.0 3337 14683555 452.2 3351 15153

A556.1 453.1 3352 15188

BHT-407-FMS-7

32If 4107ROTORCRAFT

FLIGHT MANUALSUPPLEMENT

LITTER(S)407-706-631

OR407-799-100

AND407-799-001


This supplement shall be attached to Model 407 FlightManual when LITTER(S) kit has been installed.

Information contained herein supplements informationof basic Flight Manual. For Limitations, Procedures, andPerformance Data not contained in this supplement,consult basic Flight Manual.

COPYRIGHT NOTICE Bell HelicopterCOPYRIGHT

1 9BELL® HELICOPTER TEXTRON INC.AND BELL HELICOPTER TEXTRON INC.A DIVISION OF TEXTRON CANADA LTD.ALL RIGHTS RESERVED

TEXTRONA SubsidIary of Textron Inc.


25 SEPTEMBER 1997REVISION 1 —16 SEPTEMBER 1999

BHT-407-FMS-7

NOTICE PAGE



Fort Worth, Texas 761 01-0482

NP

LOG OF REVISIONS

BHT-407-FMS-7

Original .0.14 Feb 96 RevisIon .1 .16 SEP 99Revision 1 25 Apr 96Reissue 0 09 MAY 97Reissue 0 25 SEP 97

LOG OF PAGES

CIDi/il1I2.

REVISIONNO.

REVISIONNO.

A

NOTE


Rev.1 A

PAGEFLIGHT MANUAL

PAGE

Title 1NP 0A—B 1

Original .0. 14 Feb 96Revision 1 25 Apr 96Reissue 0 09 MAY 97

CHIEF, FLIGHT TESTFOR DIRECTOR — AIRCRAFT CERTIFICATIONTRANSPORT CANADA

B Rev.1

Reissue 0 25 SEP 97Revision 1 16 SEP 99

BHT-407-FMS-7

LOG OF DOT APPROVED REVISIONS

DATE

BHT-407-FMS-7


Original 0 14 Feb 96 Reissue 0 25 SEP 97Revision 1 25 Apr 96 Revision 1 24 SEP 99Reissue 0 09 MAY 97

Rev.1 C/D

BHT-407-FMS-7

GENERAL INFORMATION

Litter kit (407-706-631) provides helicopter with capability to carry one patient on litterwith room and access for medical attendants. The principal configuration consists of twoparts, basic provisions kit and litter assembly kit. Basic provisions contain structuralbrackets and all necessary hardware for supporting a litter. Litter assembly contains afolding aluminum litter with patient restraints. In addition, an optional injured skierprovisions kit is available. In this configuration, horizontal support bar located behind co-pilot seat is moveable and may be secured with quick release pins in either normal or anupper location. This feature provides an additional 6 inches (15.24 cm) of clearance abovepatient when support bar is installed in upper position. Basic litter provisions with litterassembly adds 27 pounds (12.3 kilograms) to empty weight of helicopter. Injured skierprovisions kit adds an additional 1.5 pounds (0.7 kilograms) to empty weight.

Customized litter kit (407-799-100) is same as basic litter kit installation with injured skierprovisions, except that support bar is bolted in place which may be desired for apermanent EMS configured helicopter.

Dual Litter Kit (407-799-001) provides helicopter with capability to carry two litter patients.This kit contains structural supports and all necessary hardware to install a second litterabove standard litter. Basic litter kit provisions (407-706-631) must be installed inconjunction with this kit. If helicopter is equipped with injured skier provisions kit, duallitter kit allows upper litter patient to be placed in elevated foot position.

i/u



1-5. CONFIGURATION

Copilot cyclic and collective controls shallbe removed and stowed when litter isinstalled.

Patient(s) shall be restrained by litterstraps.


Actual weight change shall be determinedafter kit is installed and ballast readjusted,if necessary, to return empty weight CG towithin allowable limits. Refer to Center ofgravity vs weight empty chart in BHT-407-MM-i.

1-20. INSTRUMENTMARKINGS AND PLACARDS

STRUCTURAL SUPPORT MUST BEINSTALLED IN THE UPPER POSITION ORLOWER POSITION FOR FLIGHT.

Location: On copilot seat back supportassembly and on forward side of verticaltunnel.

This placard applicable only with AuxiliaryLitter Kit with Injured Skier Provisionsinstalled.

CO-PILOT SEAT SHALLNOT BE OCCUPIED UNLESSPROTECTIVE COVERS ARE

INSTALLED ON UPPERLITTER SUPPORT BRACKETS

TYPICAL

Location: On forward side of interior trimpanel centered on door post betweenupper and lower litter support brackets.

This placard applicable with basic LitterI Kit installed.

1/2

BHT-407-FMS-1 7

13211407ROTORCRAFT

FLIGHT MANUAL

SUPPLEMENTCARGO TIE-DOWNPROVISIONS KIT

407-705-201

CERTIFIED1 APRIL 1996

This supplement shall be attached to Model 407Flight Manual when CARGO TlEDOWNPROVISIONS KIT kit has been installed.


_________________ Bell Helicopterki *41L1.iiCOPYRIGHT NOTICE A Subsidiary of Textron Inc.

COPYRIGHT 1996BELL HELICOPTER INC. OFFICE BOX 482 • FORT WORTH, TEXAS 76101

AND BELL HELICOPTER TEXTRON INC.A DIVISION OF TEXTRON CANADA LTD. 1 APRIL

ALL RIGHTS RESERVED

BHT-407-FMS-1 7

NOTICE PAGE




NP

BHT-407-FMS-17

LOG OF REVISIONS

Original .0 . 1 APR 96LOG OF PAGES


FLIGHT MANUAL C/D 0i/u 0

Title—NP 0 1/2 0A—B 0

NOTE


A



Original .0 .1 APR 96

'ppAP,9Y,

DATE: / "

CHIEF, FLIGHT TESTFORDIRECTOR — AIRWORTHINESS BRANCHDEPARTMENT OF TRANSPORT

B

BUT-407-FMS-17


Original 0 1 APR 96

C/D

BHT-407-FMS-17

GENERAL INFORMATION

Cargo tie-down provisions kit (407-705-201) provides forward bulkhead tie-downprovisions using four (4) shackle/eyebolt assemblies and floor mounted provisions usingfour (4) anchor plates. These provisions allow cargo to be secured with a tie-downassembly.

i/il


Section ILIMITA TIONS

1-20. INSTRUMENTMARKINGS AND PLACARDS

Location: Each side of center post.

NO CARGO ABOVE THIS LINE

W.L. 55

Location: Each side of center post.

1/2

CAUTION

WHEN AFT FACING SEAT AREA ISUSED FOR CARGO:-DO NOT REMOVE SEAT CUSHIONS.-MAXIMUM ALLOWABLE CARGO

WEIGHT - 100 LBS.-CARGO MUST BE SECURED TO PREVENT

IN FLIGHT MOVEMENT.-CARGO WEIGHT TO BE UNIFORMLY

DISTRIBUTED.

BHT-407-FMS-20

'!407ROTORCRAFT

FLIGHT MANUAL

SUPPLEMENTKLN 89B GPS NAViGATOR

407-705-001


This supplement shall be attached to Model 407 FlightManual when KLN 89B GPS NAVIGATOR kit has beeninstalled.


Bell HelicopterkE *iit•1A Subsidiary of Textron Inc.


14 FEBRUARY 1996REVISION 1 —26 NOVEMBER 1996

COPYRIGHT NOTICECOPYRIGHT 1996

BELL HELICOPTER INC.AND BELL HELICOPTER TEXTRON INC.A DIVISION OF TEXTRON CANADA LTD.

ALL RIGHTS RESERVED

BHT-407-FMS-20

NOTICE PAGE




NP

BHT-407-FMS-20

LOG OF REVISIONS

Original .0 .14 FEB 96

Revision 1 26 NOV 96

LOG OF PAGES

REVISION REVISION

PAGE NO. PAGE NO.

FLIGHT MANUAL A — B 1

CID 1

Title 1 i/u 1

NP 0 1—2 1

NOTE


Rev.1 A



Original ........... 0 14 FEB 96Revision 1 26 NOV 96

APPROVED: DATE:

' , zCHIEF, FLIGHT TESTFORDIRECTOR — AIRCRAFT CERTIFICATION BRANCHDEPARTMENT OF TRANSPORT

B Rev.1

BHT-407-FMS-20


Original .0 .14 FEB 96Revision 1 26 NOV 96

Rev.1 C/D

GENERAL INFORMATiON

BHT-407-FMS-20

The KLN 89B GPS Navigator is anavigator's aid for use in ICAO definedworldwide geographic regions as definedin the King KLN Pilots Guide.

The system consists of a combined GPSreceiver and navigational computer, anantenna, and associated wiring. VisualNavigation data is presented on the GPS

unit. If GPS is coupled to the KCS 55Agyrocompass with Kl 525A HSI Kit 407-705-002 the system will additionally include aNAV/GPS Switch/Annunicator.

Visual navigation data, when selected, ispresented on the pilot HSI in the form of LiR steering, bearing-to-waypoint and TO!FROM indications.

Rev. 1 i/u

I


Section 1LIM1TA TIONS

1-1. INTRODUCTIONA KLN 89B Pilots Guide (King p/n 006-08786-0000, Operational Revision Status01) shall be accessible by the flight crewat all times during flight.

The GPS navigator shall be operated inaccordance with the manufacturesinstruction with the following exceptions:

1. There is no air data or fuelmanagement data available inthis installation.

2. It is the responsibility of the pilotto verify that any navigation dataused is correct.

1-20. INSTRUMELNIMARKINGS AND PLACARDS

f GPS LIMITED TO VFR USE ONLY

Lsection 2NORMAL PROCEDURES

2-3. PREFLIGHT CHECK 2-7. BEFORE TAKEOFF

2-3-A. CABIN TOP

GPS antenna — Condition and security.


2-4-A. PRESTART CHECK

GPS and CAUTION LIGHTS circuitbreakers — In.

GPU unit — Verify off.

2-7-A. GPS

GPS unit — Turn on, verify operationalrevision status on initial page is identicalto that of available KLN 89B Pilot's Guide.

Pilots HSI course pointer (if installed) —Align to desired course shown on GPSdisplay.

NAVIGPS switch-annunciator (if installed)— Press, verify GPS segment illuminatedand NAV segment extinguished.

Rev.1 1

I


NOTE

For additional normal procedures,Pilot HSI deviation bar (if installed) except air data and fuelVerify centered and TO indication management data, refer to KLNdisplayed. 89B Pilot's Guide.

Section 3EMERGENCWMALFUNCTION PROCEDURES

31. INTRODUCTION

NOTE

If GPS navigation systembecomes inoperative, continuebasic VFR navigation procedures.

2 Rev.1


SUPPLEMENT

FIRE DETECTION SYSTEM407-799-004

OR407-706-015

OR407-706-025

CERTIFIED2 MAY 1996

This supplement shall be attached to Model 407 Flight Manual when FIRE DETECTION SYSTEM has been installed.

Information contained herein supplements information in the basic Flight Manual. For Limitations, Procedures, and Performance Data not contained in this supplement, refer to the basic Flight Manual.

BHT-407-FMS-21

REISSUE — 07 JULY 2004

COPYRIGHT NOTICE

COPYRIGHT 2004



CANADA LTD.

ALL RIGHTS RESERVED

NOTICE PAGE







NP———07 JUL 2004

BHT-407-FMS-21

These data are proprietary to Bell Helicopter Textron Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation is forbidden without prior written authorization from Bell Helicopter Textron Inc.


Original ..................... 0......................02 MAY 96Reissue ..................... 0...................... 08 SEP 98

Reissue......................0 ...................... 07 JUL 04

07 JUL 2004———A/B

BHT-407-FMS-21

NOTE


LOG OF REVISIONS

LOG OF PAGES


FLIGHT MANUAL

Title.............................................................0NP ...............................................................0A/B..............................................................0C/D..............................................................0E/F ..............................................................0i/ii ................................................................01 — 2 ..........................................................0

APPROVED DATE

07 JUL 2004———C/D

BHT-407-FMS-21


Reissue......................0 ...................... 07 JUL 04


CHIEF, FLIGHT TEST FOR DIRECTOR — AIRCRAFT CERTIFICATION TRANSPORT CANADA



Reissue......................0 ......................02 SEP 04

07 JUL 2004———E/F

BHT-407-FMS-21


GENERAL INFORMATION

07 JUL 2004———i/ii

Bell Fire Detection System (407-799-004, 407-706-015 or 407-706-025) will illuminate ENGINE FIRE warning light on instrument panel if an excessive temperature or fire develops in engine compartment.


07 JUL 2004———1


11-6. WEIGHT AND CENTER OF

GRAVITY

Actual weight change shall be determinedafte r system is insta l led and ba l las t

readjusted, if necessary, to return emptyweight CG to within allowable limits.


22-4. INTERIOR AND PRESTART

CHECK

FIRE DET TEST switch — Press, ENGINE FIRElight illuminates, release, ENGINE FIRE lightextinguishes.


2———07 JUL 2004


3Table 3-1.

PANEL WORDING FAULT CONDITION CORRECTIVE ACTION

ENGINE FIRE Excessive temperature condition in engine compartment

Immediately enter autorotation.

Throttle — Close.

FUEL VALVE switch — OFF.

If time permits, FUEL BOOST/XFR circuit breaker switches — OFF

Execute a normal autorotation and landing

BATT switch — OFF.

NOTE

Do not restart engine until cause of fire has been determined and corrected.

B HT-407-FMS-22

'!407ROTORCRAFT

FLIGHT MANUAL

SUPPLEMENTAUXILIARY VERTICAL FIN

STROBE LIGHTS407-899-023

CERTIFIED10 MAY 1996

This supplement shall be attached to Model 407Flight Manual when AUXILIARY VERTICAL FINSTROBE LIGHTS have been installed.


________________ Bell Helicopterki 1 t.lflCOPYRIGHT NOTICE A Subsidiary of Textron Inc.

COPYRIGHT 1996BELL HELICOPTER INC POST OFFICE BOX 482 • FORT WORTH, TEXAS 76101

AND BELL HELICOPTER TEXTRON INC.A DIVISION OF TEXTRON CANADA LTD. 1 MA

ALL RIGHTS RESERVED

BHT-407-FMS-22

NOTICE PAGE




NP

BHT-407-FMS-22

LOG OF REVISIONS

Original .0 . 10 May 96LOG OF PAGES


Title—NP 0 i/u 0A—B 0 1—2 0CID 0

NOTE


A



Original ........... 0 . 10 May 96

APPROVED: DATE:

'7a'y 6o CHIEF, FLIGHT TESTte4FoR

DIRECTOR — AIRWORTHINESS BRANCHDEPARTMENT OF TRANSPORT

B

BHT-407-FMS-22


Original ........... 0 . 10 May 96

do

BHT-407-FMS-22

GENERAL iNFORMATION

Auxiliary vertical fin strobe lights installation (407-899-023) consist of power supply unitand two strobe lights installed on left and right auxiliary vertical fins.

i/u


LII Section 1LIMITA TIONS

1-5. CONFIGURATION

1-5-A. OPTIONAL EQUIPMENT

Auxiliary vertical fin strobe lights are notapproved for night operations.

NOTE

High intensity strobe lightsshould not be used inflight whenthere is an adverse reflection from

clouds or other weatherphenomena.

1-20. INSTRUMENTMARKiNGS AND PLACARDS

NIGHT OPERATION OF AUXILIARY VERTICALFIN STROBE LIGHTS IS PROHIBITED

(Located on inst. panel - typical)

Lsection 2NORMAL PROCEDURES

2-1. INTRODUCTION

NOTE

Both auxiliary vertical fin strobelights are controlled by AUX VERT

FIN LT on/off CCT BKR/switchlocated on overhead console.

1


Lsection 3EMERGENCWMA LFUNCTION PROCEDURES

3-7. ELECTRICAL SYSTEM strobe lights may be disabled byselecting OFF at AUX VERT FINLT CCT BKR/switch. If auxiliary

NOTE vertical fin strobe lights becomeinoperative, continue basic flight

For emergency or malfunction procedures.conditions, auxiliary vertical fin

2

B HT-407- FM 5-23

'!407ROTORCRAFT

FLIGHT MANUAL

SUPPLEMENT

RYAN TRAFFIC COLLISIONAVOIDANCE DEVICE

407-899-022CERTIFIED

15 MAY1996

This supplement shall be attached to Model 407 FlightManual when RYAN TRAFFIC COLLISION AVOiDANCEDEVICE ATS9000 has been installed in accordance with407-899-022.


________________ Bell Helicopter ki* I•1COPYRIGHT NOTICE 1 A Subsidiary of Textron Inc.

COPYRIGHT 1996BELL HELICOPTER INC I POST OFFICE BOX 482 - FORT WORTH, TEXAS 78101

AND BELL HELICOPTER TEXTRON INC.A DIVISION OF TEXTRON CANADA LTD.

ALL RIGHTS RESERVED

BHT-407-FMS-23

NOTICE PAGE


Manufacturer's Data portion of this supplement is proprietary to BellHelicopter Textron Inc. Disclosure, reproduction, or use of thesedata for any purpose other than helicopter operation is forbiddenwithout prior written authorization from Bell Helicopter Textron Inc.




NP

BHT-407-FMS-23

LOG OF REViSIONS

Original ........... 0 . 15 May 96

LOG OF PAGES

REVISION REVISION

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FLIGHT MANUAL 314 05/6 0

Title—NP 0A—B 0 MANUFACTURER'SDATACID 0i/u 0 7/8 0.1—2 0

NOTE


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LOG OF APPROVED REViSIONS

Original .0 . 15 May 96

APPROVED: DATE:

CHIEF, FLIGHT TESTFORDIRECTOR AIRWORTHINESS BRANCHDEPARTMENT OF TRANSPORT

B

BHT-407-FMS-23


Original .0 . 15 May 96

C/D

BHT-407-FMS-23

GENERAL INFORMATiON

RYAN ATS9000 Traffic and Collision Avoidance Device (TCAD) (407-899-022) consists ofdisplay unit, processor unit, transponder coupler, dual antenna module, two antennas,wiring and hardware necessary for installation. A digital display is mounted in instrumentpanel and contains all controls required to operate TCAD. Processor unit, transpondercoupler and dual antenna module are at various locations throughout helicopter,depending on configuration. Antenna locations are on cabin top and underside ofhelicopter.

RYAN TCAD is an on-board air traffic display used to identify potential collision threats.TCAD computes relative altitude and distance of threats using transponder replies fromnearby Mode C equipped aircraft. Aircraft with non-Mode C transponders can providedistance information. TCAD will not detect aircraft without operating transponders. Withincertain limits system creates a shield of airspace around helicopter, whereby detectedtraffic cannot penetrate without generating an alert. Shield size is selectable for variousphases of flight and is adjustable by pilot.

Display is a bright, alphanumeric character. Distance is displayed in nautical miles (NM)and relative altitude is displayed in 100 foot increments. TCAD is capable of displayingmultiple threats.

For familiarization of all ATS9000 TCAD features and operation, refer to Pilots Handbook,P/N 32-21 02 Revision 1 or later.

i/il




GRAVITY gravity vs weight empty chart in BHT-407-MM-i.

Actual weight change shall be determinedafter kit is installed and ballast readjusted,

L Section 2NORMAL PROCEDURES



MIJTE/PWR button — Pull (off)

MUTE/PWR button — Push (on).

1


Section 3EMERGENC V/MALFUNCTION PROCEDURES

3-7 ELECTRICAL SYSTEM 1. MUTE/PWR button — Pull (off)and continue flight.

3-7-A. TCAD MALFUNCTION

TCAD malfunction is annunciated bywords Signal Fail, SgnlFail, Link Failure orInterface Fail displayed on TCAD display.

Table 3-1.


TRAFFIC Proximate traffic detected. Locate intruder aircraft using see(advisory) and avoid concept.

NOTE

TCAD is advisory only. Operationsshall be conducted in accordancewith operational regulations ineffect at helicopter location.

2




3/4

Lsection 5

BHT-407-FMS-23

WEIGHT AND BALANCE


5/6

MANUFACTURER'S DATA BHT-407-FMS-23

L Section 1SYSTEMS DESCRIPTION

: :.,.,I TRAFFICj tt:)

Figure 1-1. Caution and waring panel

7/8


SUPPLEMENT

QUIET CRUISE MODE407-706-016

CERTIFIED8 MAY 1998

This supplement shall be attached to the BHT-407-FM-1when the Quiet Cruise Mode kit is installed.


BHT-407-FMS-25

REISSUE — 17 DECEMBER 2002REVISION 1 — 30 JULY 2008


ALL RIGHTS RESERVED

NOTICE PAGE







NP———Rev. 1—30 JUL 2008

BHT-407-FMS-25

These data are proprietary to Bell Helicopter Textron Inc. Disclosure,reproduction, or use of these data for any purpose other than helicopteroperation or maintenance is forbidden without prior written authorizationfrom Bell Helicopter Textron Inc.


Original ..................... 0......................08 MAY 98Reissue ..................... 0......................18 MAY 99

Reissue......................0 ..................... 17 DEC 02Revision.....................1 ...................... 30 JUL 08

30 JUL 2008—Rev. 1———A/B

BHT-407-FMS-25

NOTE


LOG OF REVISIONS

LOG OF PAGES


FLIGHT MANUAL

Title.............................................................1NP ...............................................................1A/B..............................................................1C/D..............................................................1E/F ..............................................................1i/ii ................................................................11 – 7............................................................08 ..................................................................19 – 12 ..........................................................013/14 ...........................................................015/16 ...........................................................0

MANUFACTURER’S DATA

17 – 18........................................................ 0


APPROVED DATE

30 JUL 2008—Rev. 1———C/D

BHT-407-FMS-25

Original ..................... 0......................08 MAY 98Reissue ..................... 0......................18 MAY 99

Reissue......................0 ..................... 17 DEC 02Revision ....................1 ...................... 30 JUL 08



Original ..................... 0....... Not FAA ApprovedReissue ..................... 0......................27 MAY 99

Reissue......................0 ......................24 SEP 03Revision ....................1 ..................... 18 AUG 08

30 JUL 2008—Rev. 1———E/F

BHT-407-FMS-25

BHT-407-FMS-25

GENERAL INFORMATION

30 JUL 2008—Rev. 1———i/ii

Installation of Quiet Cruise Mode kit (407-706-016) permits flight operations at 92% NR whenabove 50 KIAS and 200 feet AGL. Flyover noise level is reduced by 3.8 dBA SEL when in QuietCruise Mode. Kit consists of electrical selector switch on collective, annunciator on instrumentpanel, and additional markings on dual tachometer.

17 DEC 2002 1




Quiet cruise mode is approved for VFRoperations only.

1-5. CONFIGURATION


FADEC system software 5.201 or higher isrequired for Quiet Cruise Mode operations.

1-5-B. OPTIONAL EQUIPMENT

Helicopters S/N 53000 – 53074 shall be incompliance with Technical Bulletin 407-96-2(Increase in VNE).

1-5-D. CARGO HOOK

Cargo hook operations while in Quiet CruiseMode is not approved.


Actual weight change shall be determinedafter kit is installed and ballast readjusted, ifnecessary, to return empty CG to withinallowable limits. Refer to Center of gravity vsweight empty chart in BHT-407-MM-2.

1-6-A. WEIGHT

Maximum GW for Quiet Cru ise Modeoperation is 5000 pounds (2268 kilograms).

1-6-B. CENTER OF GRAVITY — QUIET CRUISE MODE OPERATION

For longitudinal CG limits refer to Grossweight longitudinal center of gravity limitschart (Figure 1-1).

For lateral CG limits refer to Gross weightlateral center of gravity limits chart (Figure 1-2).

1-7. AIRSPEED

1-7-A. QUIET CRUISE MODE

NOTERefer to Section 4, HEIGHT – VELOCITYENVELOPE.

Minimum airspeed is 50 KIAS.

VNE is 100 KIAS.

1-8. ALTITUDE

NOTERefer to Section 4, HEIGHT – VELOCITYENVELOPE.

Minimum altitude is approximately 200 feetAGL.

Maximum altitude is 6,000 feet HD.


2 17 DEC 2002

1-13. POWER PLANT

1-13-B. POWER TURBINE RPM (NP)

1-13-B-1. QUIET CRUISE MODE

1-13-D. ENGINE TORQUE

Engine torque is restricted to maximumcontinuous power (93.5%) while in QuietCruise Mode.

1-15. ROTOR

1-15-A. ROTOR RPM – POWER ON1-15-A-1. QUIET CRUISE MODE

1-20. INSTRUMENT MARKINGS AND PLACARDS

Refer to Figure 1-3 for Placards and decals.

Refer to Figure 1-4 for Instrument markings.

Minimum 91.5%Continuous operation 91.5 to 92.5%Maximum continuous 92.5%

Continuous operation 91.5 to 92.5%Maximum continuous 92.5%

17 DEC 2002 3



118 119 120 121 122 123 124 125 126 127 128 129 130


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FS 127.0

QUIET CRUISE MODE AFT LIMIT

LONGITUDINAL C.G.

M407_FMS-25__FIG_1-1_(1_OF_2).WMF


4 17 DEC 2002


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M407_FMS-25__FIG_1-1_(2_OF_2).WMF

17 DEC 2002 5



-3 -2 -1 0 1 2 3


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QUIET CRUISE MODE LIMITS

LATERAL C.G.

M407_FMS-25__FIG_1-2_(1_OF_2).WMF


6 17 DEC 2002


-60 -50 -40 -30 -20 -10 0 10 20 30 40 50 60


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23 -18

QUIET CRUISE MODE LIMITS

2268

LATERAL C.G.

M407_FMS-25__FIG_1-2_(2_OF_2).WMF

17 DEC 2002 7


Figure 1-3. Placards and decals (typical)

407 AIRSPEED LIMITATIONS – KNOTS – IAS

407FS25-1-3


0 2 4 6 8 10 12 14 16 18 2052 13745 139 132 12540 140 133 126 11935 140 135 128 120 11330 140 137 129 122 115 10825 140 138 131 124 116 109 102 9520 140 140 133 125 118 111 103 96 890 140 140 140 132 125 117 110 103 95 88

-25 140 140 140 135 130 125 119 111 104 97 89-40 137 133 128 123 118 114 110 105 101 97 93

QUIET MODE VNE 100 KIASMAXIMUM QUIET MODE ALTITUDE IS 6000 FT HD

PRESSURE ALTITUDE FT x 1000407 AIRSPEED LIMITATIONS - KIAS

OAT °C


8 Rev. 1 30 JUL 2008

Figure 1-4. Instrument Markings

NP (POWER TURBINE RPM)

Quiet Cruise Mode

91.5%

407-375-008-105 407-375-008-109

Minimum

91.5 to 92.5% Continuous operation

92.5% Maximum continuous

Normal Operations

99% Minimum

99 to 100% Continuous operation

100% Maximum continuous

NR (ROTOR RPM)

85% Minimum (power off)

85 to 107% Continuous operation (power off)

107% Maximum (power off)

407_FMS_25_0001

8

17 DEC 2002 9



22-4. INTERIOR AND

PRESTART CHECK

QUIET NORMAL mode switch — NORMAL.

QUIET and ON annunciators illuminate andextinguish with FADEC lights.

2-6. SYSTEMS CHECK

2-6-E. QUIET CRUISE MODE CHECK

NOTEIf QUIET NORMAL mode switch iscycled at less than 92% NR, increasethrottle to 100% NR to reset QUIET, ONannunciator.

1. NR — 100% RPM.2. QUIET NORMAL mode switch —

QUIET.3. QUIET and ON annunciators —

Illuminated.4. NR — 92% RPM.5. Throttle — Retard below 88% NR.

Confirm RPM warning illuminates with audio.

6. Throttle — FLY detent position.

7. QUIET NORMAL mode switch — NORMAL.

8. NR — 100% RPM.9. QUIET and ON annunciators —

Extinguished.


2-9-A. IN QUIET CRUISE MODE

Flight at altitudes above 200 AGL and atairspeeds above 50 KIAS:

QUIET NORMAL mode switch — QUIET.QUIET and ON annunciator — Illuminated.NR — 92%RPM.


2-10-B. IN QUIET CRUISE MODE

Prior to descending below 200 AGL or 50KIAS:

QUIET NORMAL mode switch — NORMAL. Monitor engine parameters.NR — 100% RPMQUIET and ON annunciators — Extinguished.


10 17 DEC 2002


33-14. QUIET CRUISE MODE

OPERATION

NOTEIn Quiet Cruise Mode, low rotor audioand RPM caution light activated at 88%NR.

If Quiet Cruise Mode fails engaged, planlanding into wind. Trans ient torqueexcursions up to 100% during landing ispermitted.

NOTELandings into winds up to 35 knotsfrom azimuths of ± 45 degrees off noseof helicopter have been demonstrated.

Use of FADEC MAN mode will immediatelydeselect Quiet Cruise Mode and reset lowrotor audio and RPM light activation point to95% NR. To insure smooth transition toFADEC MAN mode, match throttle position toNG indication.

I f e i ther QUIET or ON segment is notilluminated, return QUIET MODE switch toNORMAL position.

Table 3-1. Warning (red) lights

PANEL

WORDING FAULT CONDITION CORRECTIVE ACTION

Warning (red) lights

RPM (with low RPM audio) NR below 88%. Reduce collective and ensure throttle is in FLY detent position. Light will extinguish and audio will cease when NR increases above 88%.

NR above 88%, QUIET light not illuminated.

Return to NORMAL mode.

Table 3-2. Caution (amber) and advisory (white/green) lights

PANEL

WORDING FAULT CONDITION CORRECTIVE ACTION

HYDRAULIC SYSTEM Hydraulic pressure below limit. Exit Quiet Cruise Mode, returning to 100% NR. Verify HYD SYS switch position. Accomplish hydraulic system failure procedure.

17 DEC 2002 11



44-4. HEIGHT – VELOCITY

ENVELOPE

The Height-velocity diagram (Figure 4-1)defines conditions from which a safe landingcan be made on a smooth, level, firm surface,following an engine failure. Limitationsrespecting minimum airspeed and minimumheight above ground for Quiet Cruise Modeoperation are marked on the Height-velocitydiagram for clarity. The Height-velocitydiagram is valid only when helicopter grossweight does not exceed limits of the Altitudeversus gross weight for height-velocitydiagram (Figure 4-2)

CAUTION

IF ENGINE FAILURE OCCURS DURINGFLIGHT CONDITIONS WITHIN HEIGHT-VELOCITY DIAGRAM “AVOID AREA”,SAFE LANDING MAY NOT BEPOSSIBLE.


4-7-A. CLIMB

Reduce rate of climb data 100 feet per minutewhen operating in Quiet Cruise Mode.


12 17 DEC 2002


0 10 20 30 40 50 60 70 80 90 100 110 120


0

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ETER

S 200 FT MINIMUM HEIGHT ABOVE GROUND

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UM

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SPEE

D

AVOID

M407_FMS-25__FIG_4-1.WMF




13 14 15 16 17 18 19 20 21 22 23 24


-40 -20 0 20 40 60

OAT - °C

30 34 38 42 46 50 54


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LEVE

L

Hp -

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INTE

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AL

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OSS

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UM

OA

T LI

MIT

HEIGHT-VELOCITYAVOID AREA APPLIES

M407_FMS-25__FIG_4-2.WMF

17 DEC 2002 13/14

BHT-407-FMS-25


55-1. INTRODUCTION

This section presents loading information andinstructions necessary to ensure that flightcan be performed within approved grossweight and center of gravity limitations asdefined in Section 1.

5-3. GROSS WEIGHT CENTER OF GRAVITY


Gross weight longitudinal center of gravityand Gross weight lateral center of gravitycharts for Quiet Cruise operations are inLimitations Section 1 .

For Quiet Cruise operations maintaininglongitudinal CG within limits can be achievedby the following:

With helicopter Weight empty within envelope (BHT-407-MM-2) the helicopter will stay within Quiet Cruise limits provided both pilot and co-pilot seats are occupied. This assumes a

standard crew/passenger weight of 170 pounds (77.1 kilograms) and that fuel and payload are adjusted to stay within Maximum Gross Weight limit.If co-pilot/forward passenger seat is unoccupied then cabin payload must be adjusted to maintain flight envelope.

For helicopters operating without respect toWeight empty enve lope the p i lo t isresponsible for ensuring that when operatingin Quiet Cruise mode, helicopter weight andCG are within limits.

For Quiet Cruise operations maintainingLateral CG within limits can be achieved bythe following:

Seats should be occupied such that maximum asymmetric loading is no more than one person (170 pounds (77.1 kilograms)).

With this arrangement, a helicopter whosebasic lateral CG is ±0.3 inch (7.62 mm), willremain within lateral limits.

17 DEC 2002 15/16

17 DEC 2002 17

MANUFACTURER’S DATA BHT-407-FMS-25

Section 1SYSTEM DESCRIPTION

1This kit incorporates a two position switch oncollective (Figure 1-1) permitting pilotselection of operation at 100% NR or 92% NR(Quiet Cruise Mode). A two (2) segmentannunciator located on instrument panel(Figure 1-2), when illuminated, displaysQUIET and ON. QUIET light is illuminatedwhen low rotor audio and RPM light activationpoint is reset to 88% NR. ON l ight isilluminated when FADEC is not in 100% NRmode. Both segments are green and willdenote operat ion at 92% NR has beenselected. Dual tachometer has additionalmarkings to reflect permissible operation at92% NP.

NOTESelection of FADEC switch to MANmode, for training purposes, while inQuiet Cruise Mode, will immediatelydeselect Quiet Cruise Mode and resetlow rotor audio and RPM lightactivation point to 95% NR (triggeringlow rotor audio and RPM light). QUIETand ON segments will also extinguish.Transition back to AUTO mode shouldbe accomplished at approximately100% NP to reduce engine powertransients.

A FADEC system failure (FADEC FAILlight with FADEC fail audio), while inQuiet Cruise Mode, will retainactivation point of low rotor audio andRPM light at 88%. QUIET segment willremain illuminated and ON segmentwill extinguish. Selection of FADECMAN mode will immediately deselectQuiet Cruise Mode and reset low rotoraudio and RPM activation point to 95%NR, extinguishing the QUIET segment.

1-55. NOISE LEVELS

1-55-A. FAR PART 36 STAGE 2 NOISE LEVEL

Flyover noise level in Quiet Cruise Mode forthe Model 407 is 81.3 dBA SEL.

1-55-B. CANADIAN AIRWORTHINESS MANUAL CHAPTER 516 AND ICAO ANNEX 16 NOISE LEVEL

Flyover noise level in Quiet Cruise Mode forthe Model 407 is 81.3 dBA SEL.

BHT-407-FMS-25 MANUFACTURER’S DATA

18 17 DEC 2002

Figure 1-1. Pilot collective stick

Figure 1-2. Quiet mode annunciator

BHT-407-FMS-28


SUPPLEMENT

COPYRIGHT NOTICECOPYRIGHT BELL ® HELICOPTER TEXTRON INC. AND BELL HELICOPTER TEXTRON CANADA LTD. ALL RIGHTS RESERVED


INCREASED INTERNAL GROSS WEIGHT407-706-020


This supplement shall be attached to Model 407 FlightManual when INCREASED INTERNAL GROSS WEIGHTkit is installed.

Information contained herein supplements informationof basic Flight Manual. For Limitations, Procedures, andPerformance Data not contained in this supplement, orother applicable supplements, consult basic FlightManual.


2002

BHT-407-FMS-28

NP 16 DEC 2002




NOTICE PAGE

Manufacture’s Data portion of this supplement is proprietary to BellHelicopter Textron Inc. Disclosure, reproduction, or use of these data forany purpose other than helicopter operation is forbidden without priorwritten authorization from Bell Helicopter Textron Inc.


BHT-407-FMS-28

NOTE


LOG OF REVISIONS

Original ....................... 0................... 16 MAR 99 Reissue........................0 ................... 16 DEC 02

LOG OF PAGES


FLIGHT MANUAL

Title............................................................ 0NP.............................................................. 0A — B ........................................................ 0C/D ........................................................... 01 — 92 ..................................................... 0

16 DEC 2002 A

BHT-407-FMS-28


Original ........................0 ...................16 MAR 99 Reissue ....................... 0 ....................16 DEC 02

B 16 DEC 2002

BHT-407-FMS-28


Original ....................... 0......................7 MAY 99Reissue ....................... 0.................... 24 SEP 03

16 DEC 2002 C/D

16 DEC 2002 1



11-6. WEIGHT AND CENTER

OF GRAVITY

1-6-A. WEIGHT

Maximum approved internal gross weight fortakeoff and landing is 5250 pounds (2381k i lograms) o r as shown in IGE hoverperformance charts, Section 4.

CAUTION

LOADS THAT RESULT IN GW ABOVE5,250 POUNDS (2381 KILOGRAMS)SHALL BE CARRIED ON THE CARGOHOOK.


For longitudinal CG limits, refer to GrossWeight Longitudinal center of gravity limitscharts (Figure 1-1).

For lateral CG limits, refer to Gross WeightLateral center of gravity limits chart(Figure 1-2).

1-7. AIRSPEED

VNE is 140 KIAS, sea level to 3000 feet HD.Decrease VNE for ambient conditions inaccordance with AIRSPEED LIMITATIONSPlacards and decals (Figure 1-3).

1-8. ALTITUDE

1-8-A. DENSITY

Maximum HD for takeoff, landing, and inground effect maneuvers is 11,000 feet (3353meters).

1-8-B. DELETED


2 16 DEC 2002

Figure 1-1. Gross Weight Longitudinal center of gravity limits (sheet 1 of 2)

118 119 120 121 122 123 124 125 126 127 128 129 130


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6000

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4500

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120.5 127.6

128.7 119.8

119.0

128.0 129.0

EXTERNAL LOAD ONLY

LONGITUDINAL C.G.

M407_FMS-28__FIG_1-1_(1_OF_2).WMF

16 DEC 2002 3


Figure 1-1. Gross Weight Longitudinal center of gravity limits (sheet 2 of 2)

3000 3025 3050 3075 3100 3125 3150 3175 3200 3225 3250 3275 3300


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ILO

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AM

SLONGITUDINAL C.G.

2268

EXTERNAL LOAD ONLY

2381

3241 3061

2722

32683042

2041

1270

1202 3023

3251 3277

M407_FMS-28__FIG_1-1_(2_OF_2).WMF


4 16 DEC 2002

Figure 1-2. Gross Weight Lateral center of gravity limits (sheet 1 of 2)

-5 -4 -3 -2 -1 0 1 2 3 4 5


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-1.5

LATERAL C.G.

EXTERNAL LOAD ONLY


-0.9 1.4

1.9-1.4

2.0

-4.0 -2.5 3.0 4.0

3500

2650

5250

3500

6000

EXTERNALLOAD ONLY


EXTERNALLOAD ONLY


M407_FMS-28__FIG_1-2_(1_OF_2).WMF

16 DEC 2002 5


Figure 1-2. Gross Weight Lateral center of gravity limits (sheet 2 of 2)

-125 -100 -75 -50 -25 0 25 50 75 100 125


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EXTERNAL LOAD ONLY


EXTERNALLOAD ONLY


EXTERNALLOAD ONLY


-102 -64 76 102

1588 1588

1202

-39 52

-35 48 2268

2381

-23 36

2722

M407_FMS-28__FIG_1-2_(2_OF_2).WMF


6 16 DEC 2002

Figure 1-3. Placards and Decals (typical)

407FS28-1-3

0 2 4 6 8 10 12 14 16 18 2052 13745 139 132 12340 140 133 125 11335 140 135 128 116 10430 140 137 129 118 106 9925 140 138 131 121 109 100 93 8620 140 140 133 124 112 102 94 87 800 140 140 140 132 123 111 101 94 86 79

-25 140 140 140 135 130 125 114 102 95 88 80-40 137 133 128 123 118 114 110 105 101 93 86

PRESSURE ALTITUDE FT x 1000407 (5250 LB) AIRSPEED LIMITATIONS - KIAS

OAT °C



Location: Adjacent to existing airspeed limitations placard (typical).

16 DEC 2002 7





8 16 DEC 2002



16 DEC 2002 9



44-1. INTRODUCTION

Refer to appropriate performance charts inaccordance with optional equipment installed.

4-4. HEIGHT – VELOCITY ENVELOPE

Altitude vs gross weight for height-velocitydiagram (Figure 4-1) and Height-velocity(Figure 4-2) diagrams define conditions fromwhich a safe landing can be made on asmooth, level, firm surface following anengine failure. Height velocity diagram is validonly when helicopter gross weight does notexceed limits of the Altitude vs Gross Weightdiagram.

4-5. HOVER CEILING

NOTE

Hover performance charts are based on100% ROTOR RPM.

Satisfactory stability and control have beendemonstrated in each area of the Hoverceiling charts with winds as depicted onHover ceiling wind accountability chart (referto Basic Flight Manual).

Hover ceil ing – in ground effect charts(Figures 4-3 and 4-4) and Hover ceiling – outof ground effect charts (Figures 4-5 and 4-6)present hover performance as allowablegross weight for conditions of HP and OAT.These hovering weights are obtainable in zerowind conditions. Each chart is divided intotwo areas. Area A (non shaded area) of hover

ceiling charts presents hover performance(re la t ive to GW) for condi t ions whereadequate control margins exist for all relativewind conditions up to 35 knots, for lateral CGnot exceeding ±2.5 inches (±63 mm); and upto 17 knots, for lateral CG to ±4.0 inches (±102mm); for hover, takeoff, and landing. Area B(shaded area) of hover ceiling charts presentshover performance (relative to GW) whereadequate control margins exist for relativewinds within ± 45° of nose of helicopter up to35 knots, for lateral CG not exceeding ±2.5inches (±63 mm); and up to 17 knots, forlateral CG to ±4.0 inches (±102 mm); for hover,takeoff, and landing.


4-7-A. RATE OF CLIMB

Rate of climb (takeoff power) charts arepresented in Figure 4-7, and Rate of climb(maximum continuous power) charts arepresented in Figure 4-8.

4-10. NOISE LEVELS

4-10-A.FAR PART 36 STAGE 2 NOISE LEVEL

Model 407 is certified as a Stage 2 helicopteras prescribed in FAR Part 36, Subpart H, forgross weights up to and including certificatedmaximum takeoff and landing weight of 5250pounds (2382 kilograms).

Certified flyover noise level for Model 407 is85.5 dBA SEL.


10 16 DEC 2002

4-10-B.CANADIAN AIRWORTHINESS MANUAL CHAPTER 516 AND ICAO ANNEX 16 NOISE LEVEL

Model 407 complies with noise emissionstandards applicable to helicopter as set out

by International Civil Aviation Organization(ICAO) in Annex 16, Volume 1, Chapter 11, forgross weights up to and including certifiedmaximum takeoff and landing weight of 5250pounds (2382 kilograms).

Flyover noise level for Model 407 is 85.5 dBASEL.

16 DEC 2002 11




13 14 15 16 17 18 19 20 21 22 23 24


-40 -20 0 20 40 60

OAT - °C

30 34 38 42 46 50 54


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Hp - F

T10

,000

14,000 FT HD

MA

XIM

UM

OA

T

-200

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M O

AT

LIM

IT

12,0

00

14,0

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INT

ER

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RO

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WE

IGH

T =

525

0 L

B

DEMONSTRATEDTO 9000 FEET

DENSITY ALTITUDE

9000 FT HD

GR

OSS W

EIGH

T REG

ION

S A

B

C

D

M407_FMS-28__FIG_4-1.EMF


12 16 DEC 2002


0 10 20 30 40 50 60 70 80 90 100 110 120


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HEIGHT-VELOCITY DIAGRAMFOR GROSS WEIGHT REGIONS A TO D

0

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ID H

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ET

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NOTE: LOW HOVER POINT IS AT 6 FT SKID HEIGHT

AVOID

AV

OID

AR

EA

S

REG

ION "A" BELO

W TH

IS LINE

RE

GIO

N "B" B

ELO

W TH

IS LINE

RE

GIO

N "C

" BE

LOW

THIS

LINE

RE

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N "D

" BE

LOW

THIS

LINE

M407_FMS-28__FIG_4-2.EMF

16 DEC 2002 13



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET


M407_FMS-28__FIG_4-3_(1_OF_16).EPS


14 16 DEC 2002



ANTI-ICE ON

BASIC INLET

TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


M407_FMS-28__FIG_4-3_(2_OF_16).EPS

16 DEC 2002 15



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

BASIC INLET


M407_FMS-28__FIG_4-3_(3_OF_16).EPS


16 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET


M407_FMS-28__FIG_4-3_(4_OF_16).EPS

16 DEC 2002 17



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS





PARTICLE SEPARATOR

M407_FMS-28__FIG_4-3_(5_OF_16).EPS


18 16 DEC 2002



ANTI-ICE ON

PARTICLE SEPARATOR

TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



M407_FMS-28__FIG_4-3_(6_OF_16).EPS

16 DEC 2002 19



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

PARTICLE SEPARATOR



M407_FMS-28__FIG_4-3_(7_OF_16).EPS


20 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR



M407_FMS-28__FIG_4-3_(8_OF_16).EPS

16 DEC 2002 21



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-3_(9_OF_16).EPS


22 16 DEC 2002



ANTI-ICE ON

BASIC INLET

SNOW BAFFLES

TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


M407_FMS-28__FIG_4-3_(10_OF_16).EPS

16 DEC 2002 23



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-3_(11_OF_16).EPS


24 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-3_(12_OF_16).EPS

16 DEC 2002 25



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-3_(13_OF_16).EPS


26 16 DEC 2002



ANTI-ICE ON

PARTICLE SEPARATOR

SNOW BAFFLES

TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



M407_FMS-28__FIG_4-3_(14_OF_16).EPS

16 DEC 2002 27



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-3_(15_OF_16).EPS


28 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-3_(16_OF_16).EPS

16 DEC 2002 29




ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET


M407_FMS-28__FIG_4-4_(1_OF_16).EPS


30 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

BASIC INLET


M407_FMS-28__FIG_4-4_(2_OF_16).EPS

16 DEC 2002 31




ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

BASIC INLET


M407_FMS-28__FIG_4-4_(3_OF_16).EPS


32 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET


M407_FMS-28__FIG_4-4_(4_OF_16).EPS

16 DEC 2002 33




ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR



M407_FMS-28__FIG_4-4_(5_OF_16).EPS


34 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

PARTICLE SEPARATOR



M407_FMS-28__FIG_4-4_(6_OF_16).EPS

16 DEC 2002 35




ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

PARTICLE SEPARATOR



M407_FMS-28__FIG_4-4_(7_OF_16).EPS


36 16 DEC 2002



MAXIMUM CONTINUOUS POWER SKID HEIGHT 4 FT (1.2 METER)ENGINE RPM 100% HEATER AND ANTI-ICE ONGENERATOR 180 AMPS PARTICLE SEPARATOR


14 16 18 20 22 24 26GROSS WEIGHT - 100 KG

-40 -20 0 20 40 60OAT - °C

32 36 40 44 48 52 56 60GROSS WEIGHT - 100 LB

8000

6000

4000

2000

SEA LEVEL

Hp - F

T

10,0

00

11,000 FT HD

MA

XIM

UM

OA

T

-200

0

MIN

IMU

M O

AT

12,0

00

JET

TIS

ON

AB

LE

EX

TE

RN

AL

=600

0 L

B

MA

X G

WIN

TE

RN

AL

=525

0 L

B

100

-10

-20-30

-40

OAT - °C

M407_FMS-28__FIG_4-4_(8_OF_16).EMF

16 DEC 2002 37




ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET

SNOW BAFFLES

HOVER CEILING

IN GROUND EFFECT

M407_FMS-28__FIG_4-4_(9_OF_16).EPS


38 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-4_(10_OF_16).EPS

16 DEC 2002 39




ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-4_(11_OF_16).EPS


40 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-4_(12_OF_16).EPS

16 DEC 2002 41




ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-4_(13_OF_16).EPS


42 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-4_(14_OF_16).EPS

16 DEC 2002 43




ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-4_(15_OF_16).EPS


44 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-4_(16_OF_16).EPS

16 DEC 2002 45



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET


M407_FMS-28__FIG_4-5_(1_OF_16).EPS


46 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

BASIC INLET


M407_FMS-28__FIG_4-5_(2_OF_16).EPS

16 DEC 2002 47



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

BASIC INLET


M407_FMS-28__FIG_4-5_(3_OF_16).EPS


48 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET


M407_FMS-28__FIG_4-5_(4_OF_16).EPS

16 DEC 2002 49



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR



M407_FMS-28__FIG_4-5_(5_OF_16).EPS


50 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

PARTICLE SEPARATOR



M407_FMS-28__FIG_4-5_(6_OF_16).EPS

16 DEC 2002 51



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

PARTICLE SEPARATOR



M407_FMS-28__FIG_4-5_(7_OF_16).EPS


52 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR



M407_FMS-28__FIG_4-5_(8_OF_16).EPS

16 DEC 2002 53



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-5_(9_OF_16).EPS


54 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-5_(10_OF_16).EPS

16 DEC 2002 55



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-5_(11_OF_16).EPS


56 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-5_(12_OF_16).EPS

16 DEC 2002 57



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-5_(13_OF_16).EPS


58 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-5_(14_OF_16).EPS

16 DEC 2002 59



TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-5_(15_OF_16).EPS


60 16 DEC 2002


TAKEOFF POWER

ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-5_(16_OF_16).EPS

16 DEC 2002 61




ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET


M407_FMS-28__FIG_4-6_(1_OF_16).EPS


62 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

BASIC INLET


M407_FMS-28__FIG_4-6_(2_OF_16).EPS

16 DEC 2002 63




ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

BASIC INLET


M407_FMS-28__FIG_4-6_(3_OF_16).EPS


64 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET


M407_FMS-28__FIG_4-6_(4_OF_16).EPS

16 DEC 2002 65




ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR



M407_FMS-28__FIG_4-6_(5_OF_16).EPS


66 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

PARTICLE SEPARATOR



M407_FMS-28__FIG_4-6_(6_OF_16).EPS

16 DEC 2002 67




ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

PARTICLE SEPARATOR



M407_FMS-28__FIG_4-6_(7_OF_16).EPS


68 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR



M407_FMS-28__FIG_4-6_(8_OF_16).EPS

16 DEC 2002 69




ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-6_(9_OF_16).EPS


70 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-6_(10_OF_16).EPS

16 DEC 2002 71




ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-6_(11_OF_16).EPS


72 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS



BASIC INLET

SNOW BAFFLES


M407_FMS-28__FIG_4-6_(12_OF_16).EPS

16 DEC 2002 73




ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-6_(13_OF_16).EPS


74 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS


ANTI-ICE ON

PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-6_(14_OF_16).EPS

16 DEC 2002 75




ENGINE RPM 100%

GENERATOR 180 AMPS


HEATER ON

PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-6_(15_OF_16).EPS


76 16 DEC 2002



ENGINE RPM 100%

GENERATOR 180 AMPS



PARTICLE SEPARATOR

SNOW BAFFLES



M407_FMS-28__FIG_4-6_(16_OF_16).EPS

16 DEC 2002 77



RATE OF CLIMB





0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OAT LIM

IT

OAT-°C

30

20

10

0

-10

-20-30

-40

40

50

M407_FMS-28__FIG_4-7_(1_OF_8).WMF


78 16 DEC 2002


RATE OF CLIMB




60 KIASHEATER ON

BASIC INLET

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

20

10

0

-10

-20

-30

-40

M407_FMS-28__FIG_4-7_(2_OF_8).WMF

16 DEC 2002 79



RATE OF CLIMB





60 KIASHEATER OFF

PARTICLE SEPARATOR - PURGE OFF

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

-20

-30

-40

40

50

M407_FMS-28__FIG_4-7_(3_OF_8).WMF


80 16 DEC 2002


RATE OF CLIMB





60 KIASHEATER ON


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

20

10

0

-10

-20

-30

-40

M407_FMS-28__FIG_4-7_(4_OF_8).WMF

16 DEC 2002 81



RATE OF CLIMB





SNOW DEFLECTOR

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MAX OAT

LIMITOAT-°C

30

20

10

0

-10

-20

-30

-40

M407_FMS-28__FIG_4-7_(5_OF_8).WMF


82 16 DEC 2002


RATE OF CLIMB




60 KIASHEATER ON


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 10

00

ME

TE

RS

OAT-°C

20

10

0

-10

-20-30

-40

M407_FMS-28__FIG_4-7_(6_OF_8).WMF

16 DEC 2002 83



RATE OF CLIMB





60 KIASHEATER OFF


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MAX OAT

LIMITOAT-°C

30

20

10

0

-10

-20

-30

-40

M407_FMS-28__FIG_4-7_(7_OF_8).WMF


84 16 DEC 2002


RATE OF CLIMB





60 KIASHEATER ON


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

20

10

0

-10

-20

-30 -40

M407_FMS-28__FIG_4-7_(8_OF_8).WMF

16 DEC 2002 85



RATE OF CLIMB





0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

-20

-30

-40

40

50

M407_FMS-28__FIG_4-8_(1_OF_8).WMF


86 16 DEC 2002


RATE OF CLIMB




60 KIASHEATER ON

BASIC INLET

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

20

10

0

-10

-20

-30

-40

M407_FMS-28__FIG_4-8_(2_OF_8).WMF

16 DEC 2002 87



RATE OF CLIMB





60 KIASHEATER OFF


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

MA

X OA

T LIMIT

OAT-°C

30

20

10

0

-10

-20

-30

-40

40

50

M407_FMS-28__FIG_4-8_(3_OF_8).WMF


88 16 DEC 2002


RATE OF CLIMB





60 KIASHEATER ON


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

20

10

0

-10

-20

-30

-40

M407_FMS-28__FIG_4-8_(4_OF_8).WMF

16 DEC 2002 89



RATE OF CLIMB





SNOW DEFLECTOR

0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

30

20

10

0

-10

-20

-30

-40

M407_FMS-28__FIG_4-8_(5_OF_8).WMF


90 16 DEC 2002


RATE OF CLIMB




60 KIASHEATER ON


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

20

10

0

-10

-20

-30

-40

M407_FMS-28__FIG_4-8_(6_OF_8).WMF

16 DEC 2002 91



RATE OF CLIMB





60 KIASHEATER OFF


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

S

OAT-°C

30

20

10

0

-10

-20

-30

-40

M407_FMS-28__FIG_4-8_(7_OF_8).WMF


92 16 DEC 2002


RATE OF CLIMB





60 KIASHEATER ON


0 200 400 600 800 1000 1200 1400 1600 1800 2000


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

PR

ES

SU

RE

AL

TIT

UD

E

- F

EE

T

0 1 2 3 4 5 6 7 8 9 10


0

1

2

3

4

5

6

PR

ES

SU

RE

AL

TIT

UD

E

- 1

00

0 M

ET

ER

SOAT-°C

20

10

0

-10

-20

-30

-40

M407_FMS-28__FIG_4-8_(8_OF_8).WMF


SUPPLEMENT

INCREASED APC STARTERGENERATOR LOAD


15 JUNE 2005

This supplement shall be attached to Model 407 FlightManual when Increased APC Starter Generator Load kit hasbeen installed.


COPYRIGHT NOTICE

COPYRIGHT 2005



CANADA LTD.

ALL RIGHTS RESERVED

BHT-407-FMS-31

15 JUNE 2005

NOTICE PAGE







NP———15 JUN 2005

BHT-407-FMS-31

These data are proprietary to Bell Helicopter Textron Inc. Disclosure,reproduction, or use of these data for any purpose other than helicopteroperation is forbidden without prior written authorization from BellHelicopter Textron Inc.


Original ..................... 0...................... 15 JUN 05

15 JUN 2005———A/B

BHT-407-FMS-31

NOTE


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FLIGHT MANUAL

Title.............................................................0NP ...............................................................0A/B..............................................................0C/D..............................................................0E/F ..............................................................01 – 2 ............................................................0

APPROVED DATE

15 JUN 2005———C/D

BHT-407-FMS-31

Original ..................... 0...................... 15 JUN 05




Original ..................... 0...................... 30 JUN 05

15 JUN 2005———E/F

BHT-407-FMS-31


15 JUN 2005———1


11-12. ELECTRICAL

1-12-A. GENERATOR


Refer to Figure 1-5 for Instrument marking.Continuous operation,up to 10,000 feet HP

0 to 200 Amps

Maximum continuous upto 10,000 feet HP

200 Amps

Continuous operation,above 10,000 feet HP

0 to 180 Amps

Maximum continuousabove 10,000 feet HP

180 Amps

Transient, 2 minutes 200 to 300 Amps

Transient, 5 seconds 300 to 400 Amps


2———15 JUN 2005

Figure 1-5. Instrument Marking407FMS_31_0001

DC LOAD

FUEL PRESSURE

* P/N 407-375-007-109

0 to 200 Amps

200 Amps

180 Amps

300 Amps

400 Amps

8 PSI

8 to 25 PSI

25 PSI


Maximum


Maximum transient, 2 minutes



Maximum

Minimum



Bell HelicopterCOPYRIGHTBELL ® HELICOPTER TEXTRON INC.AND BELL HELICOPTER TEXTRON INC.A DIVISION OF TEXTRON CANADA LTD.

ALL RIGHTS RESERVED

2002

COPYRIGHT NOTICE

BHT-407-FMS-CAA

UNITED KINGDOM REGISTEREDHELICOPTERS

CAA CERTIFIED08 JANUARY 2002

08 JANUARY 2002

This supplement shall be attached to the Bell HelicopterModel 407 Flight Manual when the helicopter is registered inthe United Kingdom.

Information contained herein supplements information ofbasic Flight Manual. For Limitations, Procedures, andPerformance Data not contained in this supplement, consultbasic Flight Manual.


SUPPLEMENT

BHT-407-FMS-CAA

NP 08 JAN 2002




NOTICE PAGE

BHT-407-FMS-CAA

NOTE


LOG OF REVISIONS

Original ....................0 ....................... 08 JAN 02

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Title............................................................ 0NP.............................................................. 0

A — B........................................................ 01 — 4......................................................... 05/6 ............................................................. 0

08 JAN 2002 A

BHT-407-FMS-CAA CAA APPROVED

CIVIL AVIATION AUTHORITYSAFETY REGULATION GROUPAVIATION HOUSESOUTH AREAGATWICK AIRPORTGATWICKWEST SUSSEX RH6 OYR

APPROVED:


Original ....................0 ........................08 JAN 02

B 08 JAN 2002

CAA APPROVED BHT-407-FMS-CAA

PERFORMANCE

Section 44t

4-1. TAKEOFF DISTANCE OVER 100-FOOT OBSTACLE

The Takeoff Distance Over 100-Foot Obstaclechar t (F igure 4-1) provides takeof fperformance data. The takeoff is initiated froma stabilized 4-foot (1.2 meter) skid heighthover. Increase power smoothly to hoverpower plus 20% torque or Takeoff Power,whichever is less, and simultaneously startnosedown pitch rotation so that the aircraftaccelerates along a flight path within thetakeoff corridor defined by the Height-Velocitydiagram (Figure 4-4 in Basic Manual). As thehelicopter goes through 50 KIAS, start noseup rotation while increasing power to TakeoffPower. With the aircraft starting to climb,continue accelerating up to 65 KIAS. Enginepower limitations are imposed to precludeunsafe nosedown attitude while in the flightpath required to remain clear of criticalheight-velocity l imitations. Good pi lottechnique i s required to ach ieve thepublished takeoff performance. Wind factorsare not considered.

NOTE

Downwind takeoffs a re notrecommended because the publishedtakeoff distance performance cannotbe achieved.

The power should be applied at a ratesufficient to expedite the maneuver butnot so rapid as to overshoot the torque

value (approximately 6 seconds). Oncepower is set, it should not be furtheradjusted until the aircraft goes through50 KIAS. At this airspeed, start nose uprotation while increasing power toTakeoff Power. While starting to climb,continue accelerating up to 65 KIAS.

EXAMPLE:

What takeoff distance is required to clear a100-foot obstacle under the fol lowingconditions:

OAT 20°CHP 1500 feetGW 4500 pounds

SOLUTION:

Enter the Takeoff Distance Over 100-FootObstacle chart (Figure 4-1) at a pressurealtitude of 1500 feet, proceed horizontally tothe 20°C temperature l ine. Drop downvertically to the 4500 lb gross weight line andmove horizontally again to read a takeoffdistance of 1010 feet.

4-2. PARTIAL POWER CLIMB

Torque limited partial power rate of climbcharts are presented for an aircraft with basicinlet installed and with the heater and engineanti-ice both OFF (Figure 4-2).

The recommended best rate of climb airspeedis 60 KIAS.

08 JAN 2002 1

407-FMS-CAA.fm Page 1 Tuesday, January 4, 2005 4:40 PM


EXAMPLE:

Find the maximum rate of climb that can beattained using 65% torque under the followingconditions:

HEATER OFFENGINE ANTI-ICE OFFOAT -10°CHP 7000 feetGW 4000 pounds

SOLUTION:

Enter the appropriate gross weight chart,4000 lbs (Figure 4-2, Sheet 1 of 2). Starting ata pressure altitude of 7000 feet, proceedhorizontally to the -10°C temperature line.Drop down vertically and read a rate of climbof 1225 feet per minute.

2 08 JAN 2002



Figure 4-1. Takeoff Distance Over 100-Foot Obstacle

08 JAN 2002 3

700

800

900

1000

1100

1200

1300

DIS

TA

NC

ER

EQ

UIR

ED

TOC

LE

AR

OB

ST

AC

LE

-F

EE

T

-4000

-2000

0

2000

4000

6000

8000

PR

ES

SU

RE

AL

TITU

DE

-F

EE

T

TAKEOFF DISTANCEOVER 100 FOOT OBSTACLE

CONDITIONS:

ROTOR RPM 100%ROTATION AIRSPEED 50 KIASCLIMB AIRSPEED 65 KIASZERO WIND

TECHNIQUE:

LEVEL ACCELERATION FROM 4 FT (1.2 M) SKIDHEIGHT USING HOVER POWER + 20% TORQUE

(NOT TO EXCEED 100%) UP TO 50 KIAS ANDCLIMB AT TAKEOFF POWER THEREAFTER

4000

4500

5000

5250

GROSS WEIGHT - LBS

MAX OAT

OAT - °C

-10-20

-30-40

0

10

20

30

5040

3500 AND BELOW



Figure 4-2. Rate of Climb — Partial Power (Sheet 1 of 2)

4 08 JAN 2002

750 800 850 900 950 1000 1050 1100 1150 1200 1250 1300 1350


0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

PR

ES

SU

RE

AL

TIT

UD

E -

FE

ET

PARTIAL POWER RATE OF CLIMBRATE OF CLIMB LIMITED BY 65% TORQUE

ENGINE RPM 100%GENERATOR 180 AMPS

60 KIASHEATER / ANTI-ICE OFF

BASIC INLET

3020

100

-10

-30

-20

-40OAT

- DEG C

40

GROSS WEIGHT = 4000 LBS



Figure 4-2. Rate of Climb — Partial Power (Sheet 2 of 2)

08 JAN 2002 5/6

450 475 500 525 550 575 600 625 650 675 700 725 750


0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

PR

ES

SU

RE

AL

TIT

UD

E -

FE

ET

30

20

100-1

0

-30

-20

-40

OAT - DEG C

40

10

0 -10

-30-20

-40

PARTIAL POWER RATE OF CLIMBRATE OF CLIMB LIMITED BY 65% TORQUE

ENGINE RPM 100%GENERATOR 180 AMPS

60 KIASHEATER / ANTI-ICE OFF

BASIC INLET

GROSS WEIGHT = 5000 LBS


BHT-407-FMS-IAC AR

3q/4fl7ROTORCRAFT

FLIGHT MANUALSUPPLEMENT

INTERSTATE AVIATION COM M ITTEE —AVIATION REGISTER

COMMONWEALTH OF INDEPENDENTSTATES


This supplement shall be attached to Model 407 FlightManual when IAC-AR kit is installed and helicopter isregistered in the Commonwealth of Independent States.

Information contained herein supplements informationof basic Flight Manual. For Limitations, Procedures, andPerformance Data not contained in this supplement, orother applicable supplements, consult basic FlightManual.

COPYRIGHT NOTICE Bell HCOPYRIGHT

1999

BELL® HELICOPTER TEXTRON INC.AND BELL HELICOPTER TEXTRON INC.A DIVISION OF TEXTRON CANADA LTD.ALL RIGHTS RESERVED

TEXTRONA Subsidiary of Textron Inc.


20 MAY 1999

BHT-407-FMS-IAC AR

NOTICE PAGE


Manufacturer's Data portion of this manual is proprietary to BellHelicopter Textron Inc. Disclosure, reproduction, or use of these data forany purpose other than helicopter operation is forbidden without priorwritten authorization from Bell Helicopter Textron Inc.

Add itional copies of this publication may be obtained by contacting:Commercial Publication Distribution Center



NP

BHT-407FMS-IAC AR

LOG OF REVISIONS

Original 0 .20 MAY 99

LOG OF PAGES


Title 0NP 0A—Bi/u 01—6 07/8 09/10 0

NOTE


A

BHT-407-FMS-IAC AR DOT APPROVED

LOG OF DOT APPROVED REVISIONS

Original 0 20MAY99

APPROVED DATE /2

HIEF, FLIGHFORDIRECTOR AIRCRAFT CERTIFICATION BRANCHDEPARTMENT OF TRANSPORT

B

DOT APPROVED BHT-407-FMS-IAC AR

GENERAL INFORMATIONInterstate Aviation Committee — Aviation Register, Commonwealth of Independent States kitconsists of a metric altimeter for flight reference in metrically measured airspace environments,provisions for crew and passenger oxygen system and associated equipment for use, a VHFemergency radio communications system (located in bracketed compartment on interior side ofpilot door), and a fire detection system with aural and visual annunciators.

i/u



1-1. INTRODUCTION 1-17. FUEL AND OILRefer to Manufacturer's Data (BHT-4O7MD-1)for metric equivalency conversions of altitude CIS fuels which meet CIS specificationlimitations not covered in this supplement. GOST#10227 (TS1 and RT) are suitable for

use with this helicopter.

1-8. ALTITUDE CIS fuel TS-1 may be used at all ambienttemperatures.

NOTE CIS fuel RT usage is limited to ambientApproved oxygen equipment must be temperatures of -32° C (-25° F) or above.installed on helicopter prior tooperations above 2400 meters H with Anti-icing fuel additives are required for allpassengers, or 3000 meters H for crew fuels when ambient temperatures are below

only. 50 C (400 F).

1-10. ATCROUTING NOTE

Additives Fluid I (Ethylene GlycolHelicopter flights within CIS airspace are monoethyl ether = GOST 8313) andallowed only along routes covered by ATC Fluid IM (a mix or 50% Fluid I and 50%ground facilities using RBS mode. methyl alcohol = TU-6-10-1458) with a

concentration of 0.1 — 0.3% by volumePermits to fly along routes covered by ATC may be used for anti-icing with aboveground facilities using UVD (Russian) mode fuels.shall be obtained from ATC service.


Refer to Figure 1-3 for Placards and decals.

Refer to Figure 1-5 for Instrument markings.

1


FOR OPERATING TEMPERATURES BELOW 5°C140 F° OAT,C.I.S. RT AND TS-1 (GOST 10277)FUEL MUST CONTAIN

FLUID I (GOST 8313) OR FLUID IM (TU-6-1 0-14458) ADDITIVE.CONCENTRATION TO BE 0.1 - 0.3% BY VOLUME.

Typical

Figure 1-3. Placards and decals

2


Typical

Figure 1-5. Instrument markings

4O7IACAR-1-5

3



2-1 . INTRODUCTION 2-9. IN-FLIGHT OPERATIONSRefer to Manufacturer's Data (BHT-407-MD-1)

Use metric altimeter as reference instrumentfor metric equivalency conversions notwhen helicopter is operated in metric aircovered in this supplement.space.

2-4. INTERIOR AND PRESTART ENG ANTI ICE and PITOT HEATERswitches — ON, in visible moisture whenCHECK ambient temperature is at or below 5° C(400 F).

Oxygen system and associatedequipment — Check. NOTE

FIRE DET TEST switch — Press. At altitudes above 2400 meters H with

Verify following: passengers, or 3000 meters H withcrew only, use oxygen masks.

ENGINE FIRE light illuminates (withaudio). 2-9-A. ICING CONDITIONS

FIRE DET TEST switch — Release.

Verify following: Helicopter must exit any inadvertent entry intoicing conditions as soon as practical.

ENGINE FIRE light extinguishes (audiooff). 2-9-B. MANEUVERS

2-5. ENGINE START Do not exceed 600 per second Yaw rate(rotation) in hover.

ENG ANTI ICE switch — ON. In powered flight, do not exceed sink ratePITOT HEATER switch — ON. (vertical speed) of 600 feet per minute (3

meters per second) at airspeeds of 25 KIAS orAMPS indicator movement — Check. below.

Do not exceed bank angles of 35° withpassengers, or 60° with crew only.


In remote areas, ensure battery is beingrecharged prior to engine shutdown byincreasing electrical load (i.e. using landinglight), and monitoring AMPS indicator. Thiswill reduce possibility of having an unchargedbattery and subsequent failure to start engine.

4


213. COLD WEATHEROPERATIONS

Cover engine air intake and exhaust portswhen parked during snow (or expected snow)conditions.

Remove all snow and/or ice in vicinity ofengine air intake and exhaust ports prior toremoving covers and initiating engine start.

CAUTION

BLADE AND WINDSHIELD COVERSMAY FREEZE IN PLACE AND DAMAGEMAY RESULT DURING REMOVALPROCESS.

NOTE

Particular attention should be paidduring preflight checks to those areassubject to ice accumulation.

Engine starts are improved by keeping batteryin a warm environment prior to use.

NOTE

If possible, use an external powersource for first start of day.

One fuel boost pump can be usedduring start to conserve battery power.

Normal transmission and engine oil pressuremay be exceeded during start. Stabilizeengine at idle until minimum temperature andpressure limits are reached.

If second attempt at engine start is required(due to reaching starting limits during firstattempt) wait 3 minutes prior to next startattempt to allow battery heat to distributethrough battery and engine heat to flow intolubrication system.

Prior to embarking passengers, ground runhelicopter to warm cabin interior andminimize fogging of windows.

WARNING

PRIOR TO TAKEOFF, ENSURELANDING SKIDS ARE NOT FROZEN TOGROUND.

CAUTION

ABRUPT CONTROL INPUT CHANGESCAN CAUSE HELICOPTER TO YAWEXCESSIVELY WHEN PARKED ON ICE.

If operating in remote areas for extendedperiods, perform periodic engine restarts tomaintain gearbox oil temperatures and batterycharge.

Of extensive operations at temperatures near.4Qo C (4O0 F) are expected and starting isdifficult, consideration should be given tousing MIL-L-7808 lubricating oil in engine.Reference Allison Operation and MaintenanceManual for procedures related to changing oiltypes.

5


Section 3EMERGENC V/MALFUNCTION PROCEDURES

3-4. ENGINE FIRE DETECTION

Refer to Table 3-1 for wording, conditions, andassociated corrective action for engine firedetection system.

Table 3-1. Warning (red) lightsPANELWORDING FAULT CONDITION CORRECTIVE ACTION

ENGINE FIRE

(with audio)

Excessive temperaturecondition in enginecompartment

Immediately enter autorotation.

Throttle — Close.

FUEL VALVE switch — OFF.

If time permits, FUEL BOOST/XFRcircuit breaker switches — OFF.

Execute a normal autorotation andlandingBATT switch — OFF.

Do notNOTE

restart engine until cause of fire has been determined and corrected.



6

MANUFACTURER'S DATA BHT-407-FMS-IAC AR



7/8

MANUFACTURER'S DATA BHT-407-FMS-IAC AR


1-17. INSTRUMENT OPERATION Event Signals:

L Engine out1-17-D. FLIGHT DATA RECORDER

2. Generator failure

Operations with commercial Flight Data 3. Low rotor RPMRecorder (FDR) must use following

4. Transmission oil pressureparameters:5. Transmission oil temperature

Analog:6. Engine fire

1. Barometric altitude (from metric7. Radio ON.

altimeter)

2. Heading 1-48. EMERGENCY3. Vertical G-load COMMUNICATION SYSTEM4. DC voltage

5. Time 148A. VHF EMERGENCY RADIO

6. Engine torque VHF emergency radio and its instructions for7. Engine RPM. use are located in pilot access door

compartment.

9/10

BHT-407-MD-1

ROTORCRAFTMANUFACTURER’S DATA

NOTICE PAGE




NP———Rev. 2—31 JAN 2007

BHT-407-MD-1

These data are proprietary to Bell Helicopter Textron Inc. Disclosure,reproduction, or use of these data for any purpose other than helicopteroperation is forbidden without prior written authorization from BellHelicopter Textron Inc.


30 APR 2008—Rev. 4———A/B

BHT-407-MD-1

Original ......................0 ......................20 SEP 96Revision.....................1 ..................... 14 APR 97Revision.....................2 ......................24 JUN 97Revision.....................3 .....................05 AUG 98Revision.....................4 ..................... 20 DEC 99Revision.....................5 ......................15 FEB 00

Reissue ..................... 0 ......................09 DEC 02Revision .................... 1 ...................... 29 JUN 05Revision .................... 2 ...................... 31 JAN 07Revision .................... 3 ..................... 26 MAR 07Revision .................... 4 ......................30 APR 08

LOG OF REVISIONS

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Cover............................................................ 2Title .............................................................. 4NP................................................................. 2A/B................................................................ 4i/ii..................................................................21-1 – 1-13 ..................................................... 21-14............................................................... 41-15 – 1-19 ................................................... 21-20 – 1-23 ................................................... 41-24............................................................... 21-25............................................................... 41-26............................................................... 21-27 – 1-28 ................................................... 31-29 – 1-68 ................................................... 2

1-69 .............................................................. 41-70 – 1-76................................................... 21-77 – 1-78................................................... 41-79 – 1-106................................................. 21-107/1-108.................................................. 22-1 – 2-6....................................................... 22-7 – 2-16..................................................... 42-17/2-18 (Deleted) ..................................... 43-1/3-2.......................................................... 03-3 – 3-8....................................................... 03-9/3-10........................................................ 04-1/4-2.......................................................... 04-3 – 4-28..................................................... 0

NOTE

Revised text is indicated by a black vertical line. A revised page with only a vertical line next to the page number indicates that text has shifted or that non-technical correction(s) were made

on that page. Insert latest revision pages; dispose of superseded pages.

BHT-407-MD-1

GENERAL INFORMATION

31 JAN 2007—Rev. 2———i/ii

The Manufacturer’s Data is provided for use in conjunction with the basic Flight Manual and optional equipment supplements, as applicable. This manual contains useful information to familiarize the operator with the helicopter and its systems, to facilitate ground handling and servicing procedures, and to assist in flight planning and operations.

The Manufacturer’s Data is divided into four sections as follows:

Section 1 — SYSTEMS DESCRIPTION

Section 2 — HANDLING AND SERVICING

Section 3 — CONVERSION CHARTS AND TABLES

Section 4 — EXPANDED PERFORMANCE

MANUFACTURER’S DATA BHT-407-MD-1

31 JAN 2007—Rev. 2———1-1


1TABLE OF CONTENTS

Paragraph PageSubject Number NumberIntroduction ............................................................................................ 1-1 ........... 1-7Helicopter Description........................................................................... 1-2 ........... 1-7Principal Dimensions ............................................................................ 1-3 ........... 1-7Location References.............................................................................. 1-4 ........... 1-7

Fuselage Stations .............................................................................. 1-4-A ....... 1-7Water Lines......................................................................................... 1-4-B ....... 1-7Buttock Lines ..................................................................................... 1-4-C ....... 1-11

General Arrangement ............................................................................ 1-5 ........... 1-11Crew Compartment............................................................................ 1-5-A ....... 1-13Passenger Compartment .................................................................. 1-5-B ....... 1-14Baggage Compartment ..................................................................... 1-5-C ....... 1-14Tailboom ............................................................................................. 1-5-D ....... 1-14

Instrument Panel, Console, and Pedestal ........................................... 1-6 ........... 1-15Instrument Panel................................................................................ 1-6-A ....... 1-15Overhead Console and Pedestal ...................................................... 1-6-B ....... 1-15

Engine Instruments ............................................................................... 1-7 ........... 1-20Instrument Operation ........................................................................ 1-7-A ....... 1-20

Power-on BIT.................................................................................. 1-7-A-1 .... 1-20Commanded BIT ............................................................................ 1-7-A-2 .... 1-20Exceedance Monitoring................................................................. 1-7-A-3 .... 1-20

Check Instrument (CHECK INSTR) Light............................................. 1-8 ........... 1-21Clock ....................................................................................................... 1-9 ........... 1-21

Clock Operation ................................................................................. 1-9-A ....... 1-22Universal Time Setting .................................................................. 1-9-A-1 .... 1-22Local Time Setting ......................................................................... 1-9-A-2 .... 1-22Elapsed Time Count Up................................................................. 1-9-A-3 .... 1-22Elapsed Time Count Down............................................................ 1-9-A-4 .... 1-22Flight Time Reset ........................................................................... 1-9-A-5 .... 1-23Flight Time Alarm Set .................................................................... 1-9-A-6 .... 1-23Flight Time Alarm Display............................................................. 1-9-A-7 .... 1-23Test Mode ....................................................................................... 1-9-A-8 .... 1-23

Caution and Warning System............................................................... 1-10 ......... 1-23Power Plant ............................................................................................ 1-11 ......... 1-25Engine Controls — FADEC System ..................................................... 1-12 ......... 1-25

FADEC System................................................................................... 1-12-A ..... 1-25FADEC System — Operation ............................................................ 1-12-B ..... 1-27Power Up Mode and Built-in-test...................................................... 1-12-C ..... 1-27

BHT-407-MD-1 MANUFACTURER’S DATA

1-2———Rev. 2—31 JAN 2007



Start in Auto Mode............................................................................. 1-12-D..... 1-27Alternate Start — AUTO MODE ........................................................ 1-12-E ..... 1-31Start in Manual Mode......................................................................... 1-12-F ..... 1-31In-flight Auto Mode Operation .......................................................... 1-12-G..... 1-32FADEC System Faults ....................................................................... 1-12-H..... 1-33Category 1 — FADEC Fail/FADEC Manual — Direct Reversionto Manual System (DRTM) ................................................................ 1-12-I....... 1-34

The FADEC System Will Fail to the Manual Mode as Follows:.. 1-12-I-1 ... 1-34Category 2 — FADEC Degraded....................................................... 1-12-J...... 1-43Category 3 — FADEC Fault............................................................... 1-12-K..... 1-43Category 4 — Restart Fault............................................................... 1-12-L ..... 1-43FADEC Reversionary Governor (FADEC Software Version 5.356) 1-12-M..... 1-44Engine Overspeed Protection .......................................................... 1-12-N..... 1-44Engine Shutdown .............................................................................. 1-12-O..... 1-46HMU Manual Piston Parking Procedure .......................................... 1-12-P ..... 1-46Category 5 — Maintenance Advisory, FADEC System Faults —Engine Shutdown .............................................................................. 1-12-Q..... 1-47Checking FADEC Fault Codes.......................................................... 1-12-R..... 1-47

Engine Run Fault Codes — Procedure For Viewing................... 1-12-R-1.. 1-47FADEC Fault Codes — Procedure to Determine Last EngineRun Faults From Current Faults................................................... 1-12-R-2.. 1-48FauLt Code Charts — Use of ........................................................ 1-12-R-3.. 1-48

Clearing FADEC Fault Codes ........................................................... 1-12-S ..... 1-55Current Faults ................................................................................ 1-12-S-1 .. 1-55Last Engine Run Faults/Exceedances ......................................... 1-12-S-2 .. 1-56Accumulated Faults/ Exceedances.............................................. 1-12-S-3 .. 1-56

FADEC Training in Manual Mode ..................................................... 1-12-T ..... 1-56Cockpit Procedural Training on Ground (Engine Not Running) ... 1-12-U..... 1-56Flight Training in Manual Mode........................................................ 1-12-V ..... 1-56Simulated FADEC Failure Training (In-flight).................................. 1-12-W .... 1-57

Combined Engine Filter Assembly (CEFA) ......................................... 1-13......... 1-57Power Plant Ignition System ................................................................ 1-14......... 1-60Power Plant Temperature Measurement System ............................... 1-15......... 1-60Power Plant Compressor Bleed Air System ....................................... 1-16......... 1-60Engine Oil System ................................................................................. 1-17......... 1-61Engine Auto Relight .............................................................................. 1-18......... 1-61Auto Relight Caution Light ................................................................... 1-19......... 1-63

Start in AUTO Mode........................................................................... 1-19-A..... 1-63Engine Out in Auto Mode.................................................................. 1-19-B..... 1-63Start and Continuous Operation in Manual Mode .......................... 1-19-C..... 1-64

Engine ANTI ICE Switch........................................................................ 1-20......... 1-64Engine Indicators................................................................................... 1-21......... 1-64

Torque Gauge .................................................................................... 1-21-A..... 1-64Torque Gauge — Range Marking ................................................. 1-21-A-1.. 1-64Engine Torque Exceedance.......................................................... 1-21-A-2.. 1-65


31 JAN 2007—Rev. 2———1-3



Gas Producer Gauge (NG) ................................................................. 1-21-B ..... 1-66Dual Tach Gauge................................................................................ 1-21-C ..... 1-66Measured Gas Temperature Gauge ................................................. 1-21-D ..... 1-66Engine Oil Temperature/Pressure Gauge........................................ 1-21-E ..... 1-68

Engine Out Warning Light and Horn.................................................... 1-22 ......... 1-68FADEC in Auto Mode......................................................................... 1-22-A ..... 1-68FADEC in Manual Mode..................................................................... 1-22-B ..... 1-68

Engine Overspeed Warning Light ........................................................ 1-23 ......... 1-68Engine Chip Caution Light.................................................................... 1-24 ......... 1-69Fuel System Description ....................................................................... 1-25 ......... 1-69

Fuel System Operation...................................................................... 1-25-A ..... 1-69Left Fuel Boost/XFR Alternate Electrical Circuit............................. 1-25-B ..... 1-73Fuel System Controls ........................................................................ 1-25-C ..... 1-73Fuel Valve Switch............................................................................... 1-25-D ..... 1-73Left and Right Fuel Boost/XFR Switch ............................................ 1-25-E ..... 1-73Fuel Cell Drain Switches ................................................................... 1-25-F...... 1-74Fuel System Indicators...................................................................... 1-25-G..... 1-74Fuel System Capacity........................................................................ 1-25-H ..... 1-74Fuel Quantity Gauging System (FQGS) ........................................... 1-25-I....... 1-74Fuel Quantity Signal Conditioner ..................................................... 1-25-J...... 1-74Signal Conditioner Built-in-test (BIT) ............................................... 1-25-K ..... 1-76

Power-up BIT.................................................................................. 1-25-K-1 .. 1-76Continuous BIT .............................................................................. 1-25-K-2 .. 1-76

Fuel Quantity Calculation.................................................................. 1-25-L...... 1-76Fuel Quantity Gauge.......................................................................... 1-25-M..... 1-76Fuel Quantity Button ......................................................................... 1-25-N ..... 1-77Fuel Pressure/Ammeter Gauge ........................................................ 1-25-O..... 1-77Fuel Valve Light ................................................................................. 1-25-P ..... 1-77

L/Fuel and R/Fuel Boost Lights .................................................... 1-25-P-1 .. 1-77L/Fuel XFR and R/Fuel XFR Lights ............................................... 1-25-P-2 .. 1-77

Fuel Filter Light .................................................................................. 1-25-Q..... 1-78Fuel Low Light.................................................................................... 1-25-R ..... 1-79

Retirement Index Number (RIN)............................................................ 1-26 ......... 1-79Transmission.......................................................................................... 1-27 ......... 1-79

Freewheel Assembly ......................................................................... 1-27-A ..... 1-79Transmission Oil System .................................................................. 1-27-B ..... 1-81Transmission Indicators ................................................................... 1-27-C ..... 1-81Transmission Oil Temperature and Pressure Gauge ..................... 1-27-D ..... 1-83

Transmission Oil Temperature Gauge......................................... 1-27-D-1 .. 1-83Transmission Oil Pressure Light.................................................. 1-27-D-2 .. 1-83

Transmission Chip Annunciator ...................................................... 1-27-E ..... 1-83Rotor System.......................................................................................... 1-28 ......... 1-83

Main Rotor Hub and Blades .............................................................. 1-28-A ..... 1-83Tail Rotor Hub and Blades ................................................................ 1-28-B ..... 1-83Tail Rotor Gearbox............................................................................. 1-28-C ..... 1-86


1-4———Rev. 2—31 JAN 2007



Tail Rotor Chip Light ......................................................................... 1-28-D..... 1-86Rotor System Indicators ....................................................................... 1-29......... 1-86

Dual Tach Gauge ............................................................................... 1-29-A..... 1-86RPM Light and Warning Horn........................................................... 1-29-B..... 1-86

NR Less than 95% .......................................................................... 1-29-B-1.. 1-86NR 107% or Greater........................................................................ 1-29-B-2.. 1-86

Flight Control System ........................................................................... 1-30......... 1-87Rotor Controls ................................................................................... 1-30-A..... 1-87Main Rotor .......................................................................................... 1-30-B..... 1-87

Cyclic .............................................................................................. 1-30-B-1.. 1-87Collective........................................................................................ 1-30-B-2.. 1-87

Tail Rotor ............................................................................................ 1-30-C..... 1-89Airspeed Actuated Pedal Stop System............................................ 1-30-D..... 1-89

Hydraulic System................................................................................... 1-31......... 1-89Hydraulic Indicators .......................................................................... 1-31-A..... 1-90Hydraulic Filter Indicators ................................................................ 1-31-B..... 1-90Hydraulic System Light..................................................................... 1-31-C..... 1-90

Electrical System................................................................................... 1-32......... 1-90External Power................................................................................... 1-32-A..... 1-94Battery Switch.................................................................................... 1-32-B..... 1-94Battery Charging (17 and 28 Amp/Hour Batteries) ......................... 1-32-C..... 1-95Generator Switch ............................................................................... 1-32-D..... 1-95Start Switch ........................................................................................ 1-32-E ..... 1-96Electrical System Indicators............................................................. 1-32-F ..... 1-96Fuel Pressure/DC Ammeter .............................................................. 1-32-G..... 1-96

Ammeter — 200 Amps Max Scale ................................................ 1-32-G-1.. 1-96Ammeter — 400 Amps Max Scale ................................................ 1-32-G-2.. 1-96

Voltmeter ............................................................................................ 1-32-H..... 1-96Battery Hot Annunciator ................................................................... 1-32-I....... 1-97Battery Relay Annunciator................................................................ 1-32-J...... 1-97Start Annunciator .............................................................................. 1-32-K..... 1-97Gen Fail Annunciator ........................................................................ 1-32-L ..... 1-97

Pitot Static System ................................................................................ 1-33......... 1-97Basic Flight Instruments....................................................................... 1-34......... 1-97

Airspeed Indicator ............................................................................. 1-34-A..... 1-97Altimeter ............................................................................................. 1-34-B..... 1-99Vertical Speed Indicator (VSI)........................................................... 1-34-C..... 1-99Inclinometer ....................................................................................... 1-34-D..... 1-99Optional Flight Instruments.............................................................. 1-34-E ..... 1-99Attitude Indicator ............................................................................... 1-34-F ..... 1-99Directional Gyro................................................................................. 1-34-G..... 1-99Turn-and-Slip Indicator ..................................................................... 1-34-H..... 1-99Encoding Altimeter............................................................................ 1-34-I....... 1-99

Navigation Systems and Instruments.................................................. 1-35......... 1-99Magnetic Compass ............................................................................ 1-35-A..... 1-99


31 JAN 2007—Rev. 2———1-5



Optional Navigation Equipment ....................................................... 1-35-B ..... 1-100Horizontal Situation Indicator (HSI).............................................. 1-35-B-1 .. 1-100Global Positioning System (GPS) ................................................ 1-35-B-2 .. 1-100VOR Indicator ................................................................................. 1-35-B-3 .. 1-100Automatic Direction Finder (ADF) ................................................ 1-35-B-4 .. 1-101ATC Transponder........................................................................... 1-35-B-5 .. 1-101VHF NAV/COMM System ............................................................... 1-35-B-6 .. 1-101

Three or Five Place Intercommunication System............................... 1-36 ......... 1-102Normal Mode ...................................................................................... 1-36-A ..... 1-103Isolate Mode ....................................................................................... 1-36-B ..... 1-103Private Mode....................................................................................... 1-36-C ..... 1-103Aft Cabin Mode................................................................................... 1-36-D ..... 1-104

Emergency Communication System.................................................... 1-37 ......... 1-104Avionics Master Switch......................................................................... 1-38 ......... 1-104Miscellaneous Instruments................................................................... 1-39 ......... 1-104

Outside Air Temperature Indicator................................................... 1-39-A ..... 1-104Hourmeter........................................................................................... 1-39-B ..... 1-104Ventilation and Defog System .......................................................... 1-39-C ..... 1-104Lighting System ................................................................................. 1-39-D ..... 1-105Cockpit Utility Light ........................................................................... 1-39-E ..... 1-105Instrument Panel and Associated Lighting ..................................... 1-39-F...... 1-105Aft Cabin Lighting.............................................................................. 1-39-G..... 1-105Landing Lights ................................................................................... 1-39-H ..... 1-105Position Lights ................................................................................... 1-39-I....... 1-106Anticollision Light.............................................................................. 1-39-J...... 1-106

Emergency Equipment .......................................................................... 1-40 ......... 1-106Portable Fire Extinguisher .................................................................... 1-41 ......... 1-106First Aid Kit............................................................................................. 1-42 ......... 1-106Pointer 4000 ELT.................................................................................... 1-43 ......... 1-106

FIGURES

Figure PageSubject Number Number

Principal Dimensions ............................................................................ 1-1 ........... 1-8Aft Cabin and Baggage Compartment ................................................. 1-2 ........... 1-9Fuselage Assembly ............................................................................... 1-3 ........... 1-12Instrument Panel and Pedestal............................................................. 1-4 ........... 1-16Overhead Console ................................................................................. 1-5 ........... 1-18Maintenance Ports ................................................................................. 1-6 ........... 1-19Caution and Warning Panel .................................................................. 1-7 ........... 1-24Power Plant ............................................................................................ 1-8 ........... 1-26FADEC Control System Schematic ...................................................... 1-9 ........... 1-28


1-6———Rev. 2—31 JAN 2007

FIGURES (CONT)


Auto to Manual Transition at Low Fuel Flow ...................................... 1-10......... 1-36Auto to Manual Transition at Intermediate Fuel Flow ........................ 1-11......... 1-38Auto to Manual Transition at High Fuel Flow...................................... 1-12......... 1-40Power Plant Components ..................................................................... 1-13......... 1-58Engine Oil System ................................................................................. 1-14......... 1-62Fuel System............................................................................................ 1-15......... 1-70Fuel Transfer System Schematic ......................................................... 1-16......... 1-71Left Fuel Boost/XFR Alternate Circuit Schematic .............................. 1-17......... 1-75Transmission Assembly ....................................................................... 1-18......... 1-80Transmission Oil System...................................................................... 1-19......... 1-82Main Rotor Assembly ............................................................................ 1-20......... 1-84Tail Rotor Assembly .............................................................................. 1-21......... 1-85Flight Controls ....................................................................................... 1-22......... 1-88Airspeed Actuated Pedal Stop System................................................ 1-23......... 1-91Hydraulic System................................................................................... 1-24......... 1-92DC Electrical System............................................................................. 1-25......... 1-93Pitot Static System ................................................................................ 1-26......... 1-98

TABLES


Time to Power Change.......................................................................... 1-1........... 1-42250-C47B FADEC Software Version 5.202 Fault Code Display ......... 1-2........... 1-49250-C47B FADEC Software Version 5.356 (Reversionary Governor)Fault Code Display ................................................................................ 1-3........... 1-52Engine Torque Exceedance Monitoring .............................................. 1-4........... 1-65GAS PRODUCER GAUGE (NG) Exceedance Monitoring.................... 1-5........... 1-66MGT Exceedance Monitoring — During Start(P/N 407-375-001-101/103)..................................................................... 1-6........... 1-67MGT Exceedance Monitoring — During Normal Operation(P/N 407-375-001-101/103)..................................................................... 1-7........... 1-67MGT Exceedance Monitoring — During Start(P/N 407-375-001-105 and Subsequent)............................................... 1-8........... 1-67MGT Exceedance Monitoring — During Normal Operation(P/N 407-375-001-105 and Subsequent)............................................... 1-9........... 1-67Transfer Light Activation Table, S/N 53175, and Subsequent........... 1-10......... 1-78


31 JAN 2007—Rev. 2———1-7


11-1. INTRODUCTION

Bell Model 407 helicopter, primary andauxiliary systems, and emergency equipmentare described within this section. Optionalequipment systems which do not require aFlight Manual Supplement (FMS) will bedescribed in this section.

1-2. HELICOPTER DESCRIPTION

The Model 407 is a single engine, seven-placelight helicopter. Standard configurationprovides for one pilot and six passengers.

The fuselage consists of three main sections:the Forward Section, the IntermediateSection, and the Tailboom Section. Theforward sect ion u t i l i zes a luminumhoneycomb and carbon graphite structureand prov ides the major load carry inge lements o f the forward cab in . Theintermediate section is a semi-monocoquestructure which uses bulkheads, longeronsand carbon fibre composite side skins. Theta i lboom is an a luminium monocoqueconstruction which transmits all stressesthrough its external skins.

The helicopter is powered by a Rolls-Royce,Model 250-C47B engine. Refer to paragraph1-11, Power Plant, for complete systemdescription.

The main rotor is a four-bladed, soft-in-planedesign with a composite hub and individuallyinterchangeable blades. The tail rotor is atwo-bladed teetering rotor that providesdirectional control. Refer to paragraph 1-28,

Rotor System, for a complete systemdescription.

Basic helicopter landing gear is the low skidtype. Optional pop-out emergency flotationgear or high skid gear is also available.

1-3. PRINCIPAL DIMENSIONS

Principal exterior dimensions are shown inFigure 1-1. All height dimensions must beconsidered approximate due to variations inloading and landing gear deflection. Principalinterior dimensions of the cabin and baggagecompartment are shown in Figure 1-2.

1-4. LOCATION REFERENCES

Locations on and within the helicopter can bedetermined in re lat ion to the fuselagestat ions, waterlines, and buttock l inesmeasured in inches (mm) f rom knownreference points.

1-4-A. FUSELAGE STATIONS

Fuselage stations (FS or STA) are verticalplanes perpendicular to, and measured along,the longitudinal axis of the helicopter. Stationzero is the reference datum plane and is 1.0inch (25.4 mm) forward of the nose of thehelicopter or 55.16 inches (1401 mm) forwardof the forward jack point center line.

1-4-B. WATERLINES

Waterl ines (WL) are hor izontal p lanesperpendicular to, and measured along, thevertical axis of the helicopter. Waterline zerois a reference plane located 20.0 inches (508mm) below the lowest point on the fuselage.


1-8———Rev. 2—31 JAN 2007

Figure 1-1. Principal Dimensions407_MD_01_0003

1 When high gear is installed, dimension is 10.91 feet (3.33 m).

TYPICAL

NOTE


31 JAN 2007—Rev. 2———1-9

Figure 1-2. Aft Cabin and Baggage Compartment (Sheet 1 of 2)407_MD_01_0004

1.

2.

Aft cabin space 85 cubic feet (2.4 m3)

Left forward cabin 20 cubic feet (0.6 m3)

Cabin floor stressed for 75 pounds per square foot (3.7 kg/100 cm2).

Baggage compartment 16 cubic feet (0.45 m3)

Baggage compartment floor stressed for 86 pounds per square foot (4.2 kg/100 cm2).

37.0 IN.

(940 mm)

42.0 IN.

(1067 mm)

16.5 IN.

(419 mm)

21.5 IN.

(546.1 mm)

36.0 IN.

(914.4 mm) 28.0 IN.

(711.2 mm)

AUX

FUEL

TANK

20.75 IN.

(527.05 mm)

TYPICAL

BAGGAGE COMPARTMENT FLOOR

VIEW LOOKING DOWN

WITH AUX TANK INSTALLED

VIEW FROM LEFT-HAND SIDE

NOTES


1-10———Rev. 2—31 JAN 2007

Figure 1-2. Aft Cabin and Baggage Compartment (Sheet 2 of 2)407_MD_01_0005

1.

2.

3.

4.

5.

6.

1

12

11

109

8

7

6

5

2

3

4

39.3 inches (997 mm)

11.0 inches (279 mm) deep

36.5 inches (927 mm) forward side

12.0 inches (305 mm) high forward side


46.0 inches (1168 mm)

7.

8.

9.

10.

11.

12.

49.0 inches (1245 mm)


43.5 inches (1105 mm)


58.5 inches (1486 mm) including litter door


TYPICAL

VIEW FROM LEFT-HAND SIDE


31 JAN 2007—Rev. 2———1-11

1-4-C. BUTTOCK LINES

Buttock l ines (BL) are vert ica l p lanesperpendicular to, and measured to the left andright, along the lateral axis of the helicopter.Buttock line zero is a plane at the lateralcenterline of the helicopter.

1-5. GENERAL ARRANGEMENT

The fuselage assembly (Figure 1-3) consistsof three main sections: the forward sectionwhich extends from the cabin nose to thebulkhead aft of the passenger compartment,the intermediate section which extends fromthe bu lkhead a ft o f the passengercompartment to the tailboom attachmentbulkhead, and the tailboom section.

The forward section provides for pilot andpassenger seating, the fuel cell enclosures,and the pylon support. The basic structure ofthe forward fuselage utilizes two honeycombpanels which are connected to create thefloor, and an aluminum roof beam assemblywhich is attached to the center of anotherhoneycomb panel to create the roof. Twobulkhead assemblies and a center postconnect the f loor and roof to make anintegrated structure. The forward fuselagesection is closed out by two carbon fibrecompos i te s ide body fa i r ings w i thinterchangeable composite doors withautomotive type latches.

The landing gear is attached to the bottom ofthe forward and aft bulkheads. The gear usesa three point attachment configuration toprevent ground resonance. The skid typelanding gear consists of two skids attached tothe ends of two arched crosstubes that aresecured to the fuselage by means of a threepoint attachment configuration. Each skidtube is fitted with a tow fitting, two saddleswith sockets for crosstubes, skid shoes along

the bottom, a rear cap, and two eyeboltfittings for mounting of ground handling gear.

The in termed ia te sect ion is asemi-monocoque structure which usesbulkheads, longerons, and carbon fibrecomposite side skins which do not needstiffeners or stringers. The lower section ofthe intermediate fuselage is closed out by ahoneycomb composi te fa i r ing . Theintermediate section provides a deck forengine installation, a baggage compartment,and an equipment compartment under theengine deck for electrical equipment, airconditioning equipment, and the tail rotorcontrol servo actuator. The engine pan andforward and aft f irewalls of the enginecompartment are manufactured from titanium.

The tailboom is a monocoque constructionwhich transmits all stresses through itsexternal skins. To increase the strength of thestructure, certain skins incorporate bondeddoublers. The tailboom assembly includes thetail rotor driveshaft, pitch change mechanism,tail rotor gearbox, tail rotor hub and bladeassembly, and the horizontal stabilizer andvertical fin.

Cowlings and fairings which enclose thevar ious roof and ta i lboom mountedassemblies provide inspection access via theuse of hinge mounts, access doors, andinspection windows or cutouts. They aremanufactured from either composite ora luminum mater ia ls and are readi lyremovable for maintenance access.

The forward cowl ing is a composi teconstruction and incorporates a single pointhinge fitting which is located at the forwardend of the fairing. Two flush mounted togglehook latches secure the cowling. Supportrods are internally stowed on each side of thecowl and f it into roof mounted clips tosupport the cowl in a raised position.


1-12———Rev. 2—31 JAN 2007

Figure 1-3. Fuselage Assembly407_MD_01_0006

Windshield

Skylight window

Forward fairing

Transmission fairing assembly

Engine cowl

Aft fairing

Engine oil tank access door

Tail rotor driveshaft cover

Tail rotor gearbox fairing

Vertical fin

Tail skid and weight

Tailboom

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

11.

12.

Finlet

Horizontal stabilizer

Passenger door

Litter door

Crew door

Side body fairing

Oil cooler blower inlet duct

Aft skin panel

Air inlet cowl assembly

Baggage compartment

Forward fuselage

Lower window

Battery compartment door

13.

14.

15.

16.

17.

18.

19.

20.

21.

22.

23.

24.

25.

10

9

8

14

13

12

18

17

16

15

24

25

1

2

3

4

23

21

7

5

6

19

22

20

11


31 JAN 2007—Rev. 2———1-13

The transmission cowling is a compositeconstruction and encloses the forward half ofthe main transmission. This cowling ismounted with fasteners. The fasteners arecontained in metal ejector clips which ensurethey are maintained with the cowling duringremoval, and which ensure easy identificationof any fastener that may have dislodged fromits receptacle.

The air induction cowling is an aluminumconstruction and encloses the aft half of themain transmission. Inlet ducts are providedon each side of the cowling to direct theairflow into the induction screen or particleseparator. A small removable window isprovided on either side of the cowling to allowchecking of the inlet area when the particleseparator is installed. An inspection cutout isprovided on the right side to allow viewing ofthe transmission oil level sight gauge. Twohinged doors with flush-type latches are alsoprov ided for inspect ion o f the a fttransmission area.

The engine cowling has composite hingedside doors and an aluminum upper structure.The side doors incorporate flush-type latchesand wing style stud fasteners. Attached to theside doors are oil cooler blower inlet ducts(S/N 53519 and subsequent or Post ASB407-02-54). The doors are held in the openposition with mechanical support devices.The side doors and upper structure of theengine cowling incorporate screened vents toallow air movement through the enginecompartment.

The aft cowling is a composite constructionand encloses the oil cooler and blowerassembly and the engine o i l tank . I tincorporates a cutout to view the engine oiltank sight gauge, and two doors which arefastened with wing style stud fasteners. Thecowling incorporates screened cooling vents.

The ta i lboom ut i l i zes two composi teconstructed fairings to enclose the tail rotordriveshaft.

The tail rotor gearbox fairings are bothcomposite constructed and secured withscrews. The upper and lower fairings bothincorporate hinge mounted doors for access.

1-5-A. CREW COMPARTMENT

The crew compartment or cockpit occupiesthe forward part of the cabin (Figure 1-3). Thepilot station is on the right side and thecopilot/forward passenger station is on theleft.

An instrument panel is mounted on a centralpedestal in front of crew seats. The panel istilted upward for maximum visibility fromeither seat.

The overhead console is centered on theforward cabin ceiling and incorporates mostof the electrical systems circuit breakers andswitches.

Each c rew sea t is covered w i thflame-retardant fabric and is equipped with alap seat belt and a dual shoulder harness.Each shoulder harness contains an inertiareel which locks in the event of a rapiddeceleration.

A door on either side permits direct access tothe crew compartment. Each door is equippedwith interior and exterior handles whichoperate a positive bolt latching mechanism.This style of door latching mechanismensures smooth and positive operation whenopening and closing the doors. Each exteriordoor handle incorporates a lock. The doorwindows are made of gray tinted acrylicplastic and incorporate a lower forwardsliding window for ventilation.

The main windshields, lower cabin nose chinbubbles, and upper cabin roof skylightwindows are also made of gray tinted acrylicplastic.


1-14———Rev. 4—30 APR 2008

1-5-B. PASSENGER COMPARTMENT

The aft area of the cabin (Figure 1-2) containsa space of 85 cubic feet (2.4 m3) for thecarrying of passengers or internal cargo. Thecabin can be configured with utility, standard,or corporate interior kits. Each kit will include,bu t is no t l imi ted to , vacuum fo rmedpolycarbonate panels, armrests, door andsidewall magazine pockets, assist steps, doorpulls, carpet, and all trim and attaching tracksand hardware.

Basic configuration includes two aft facingand three forward facing seats. All seats arecovered with flame-retardant fabric and areequipped with lap seat belts and shoulderharnesses. The shoulder harnesses lock inthe event of a rapid deceleration.

A cargo restraint kit is available for theinstallation of tie-down provisions. This kitprov ides forward bu lkhead t ie -downprovisions using four shackle/eyeboltassemblies and floor mounted provisionsusing four anchor plates. These provisionsallow cargo to be secured with a tie-downassembly. It is the responsibility of the pilot toensure the cargo is adequately secured anduniformly distributed.

A door on either side permits direct access tothe passenger compartment. In addition, theleft passenger door is hinged on the litterdoor, so that the two may be opened togetherto aid in the loading of the helicopter. TheLITTER DOOR caution light will illuminate ifthe litter door is not properly secured. Eachdoor is equipped with interior and exteriorhandles, which operate a posit ive boltlatching mechanism. This style of doorlatching mechanism ensures smooth andpositive operation when opening and closingthe doors . Each exter ior door handleincorporates a lock. All cabin windows are

made of t inted acry l ic p last ic and thepassenger doors can incorporate a lowerforward sliding window for ventilation.

1-5-C. BAGGAGE COMPARTMENT

The baggage compartment (Figure 1-2) islocated aft of the passenger compartment andhas a capacity of 16 cubic feet (0.45 m3). Thecompartment can carry up to 250 pounds(113.4 kg) of baggage or other cargo whichcan be secured using a tie-down assemblyand the tie-down fittings provided. It is theresponsibility of the pilot to ensure the cargois adequate ly secured and uni formlydistributed.

Access to the compartment is provided by anexterior door on the left side of the aftfuselage. The door is hinged at the forwardend and opens the full width and height of thecompar tment . I t is secured by twopush-button latches and a keyed lock. TheBAGGAGE DOOR caution light will illuminateif the door is not properly secured.

A 19.2 US gallon (72.7 L) fuel tank may also beinstalled in the baggage compartment as anoptional kit (BHT-407-FMS-6).

1-5-D. TAILBOOM

The tailboom consists of the tailboom and thecomponents it supports: tail rotor and drivesystem, vertical fin, and horizontal stabilizer(Figure 1-3).

The tail rotor drive system consists of adriveshaft mounted along the top of thetailboom and a gearbox, which reduces thetail rotor driveshaft input speed from 6317 to2500 RPM. The gearbox also changes thedirection of drive 90° to accommodate the tailrotor for directional control.

31 JAN 2007 Rev. 2 1-15


The vert ical f in, composed primarily ofaluminum and honeycomb construction,provides directional (yaw) stability and ismounted on the aft end of the tailboom on theright side. It contains a top fairing to mount theanticollision light and on the lower edge, arubber bumper, and tail skid to protect the tailrotor and fin in the event of a tail low landing.The fin sweeps back, both above and belowthe tailboom. The leading edge is cantedoutboard 9° to reduce the required amount oftail rotor thrust during forward flight at cruisespeed.

The horizontal stabilizer extends through thetailboom, has leading edge slats, and smallauxiliary vertical fins or “dynamic dihedrals”.The main body of the horizontal stabilizer is aone-piece aluminum honeycomb structure. Itis an inverted a irfo i l which provides adownward resultant lift on the tailboom tomaintain the cabin in a nearly level attitudethroughout al l cruise airspeeds and toaerodynamically streamline the fuselage toreduce drag. The leading edge slats aredesigned to improve pitch stability duringclimbs.

The small auxiliary vertical fins are located oneach end of the horizontal stabilizer. Theleading edges of the fins are both offset 5°outboard of the helicopter centerline. Thisimproves the roll stability of the helicopter inforward flight.

The ta i lboom u t i l izes two composi teconstructed fairings to enclose the tail rotordriveshaft. The cowlings are mounted withstuds and receptac les . The s tuds arecontained in metal ejector clips, which ensurethe studs are maintained with the cowlingsduring removal, and which ensure the easyidentification of any stud that may havedislodged from its receptacle.

The tail rotor gearbox fairings are bothcomposite constructed and secured withscrews. A hinged door with wing style studfasteners is incorporated on the top fairing forease of servicing the gearbox. A hinged door

is incorporated on the bottom of the fairing forease of access to the tail rotor gearbox chipdetector. The tail rotor gearbox oil level maybe viewed through the aft screened coolingvent of the fairing.

1-6. INSTRUMENT PANEL,CONSOLE, AND PEDESTAL

1-6-A. INSTRUMENT PANEL

Figure 1-4 shows the instrument panel withoptional equipment installed. The instrumentpanel provides vibration dampening and ishinge mounted on a central pedestal in front ofthe crew seats.

The instrument panel is tilted at a 5° angle formaximum visibility. The flight instruments arelocated on the right side of the panel, and thesystems instruments are in two rows to the leftof the flight instruments. The caution andwarning panel is mounted just below theglareshield across the top of the instrumentpanel.

1-6-B. OVERHEAD CONSOLE ANDPEDESTAL

The overhead console (Figure 1-5) is centeredon the cabin ceiling and incorporates most ofthe electrical systems’ circuit breakers andswitches. The forward section of the panel hasintegral lighting controlled by the instrumentlighting rheostat knob.

The pedestal (Figure 1-4) extends from theinstrument panel downward and aft to thecabin seat structure. This forms a mountingplatform for optional equipment such as theaudio panel, radios, and gyrocompass controlpanel etc. The lower left side of the pedestal(F igure 1-6 ) conta ins the FADEC/ECUmaintenance bu t ton , FADEC/ECUmaintenance port, ENGINE INSTRUMENTmaintenance port, 28 VDC auxiliary receptacle,and the M/R RADS maintenance port and GPSdata loader port.

1-16 Rev. 2 31 JAN 2007


Figure 1-4. Instrument Panel and Pedestal (Sheet 1 of 2)

1.

2.

3.

4.

5.

6.

7.

8.

PSI

FUEL

1

0

2

20

AMPS

X 10

3

4

0

8

12

16

3

1

0

2

X100

QTY LBS

4 5

11

10

12

9

8

7

6

12

4

2

6

8

10

~C

X 10

0

PSI

0

5

15

10

10

5

15

20

~C

X 10

0

PSI

0

5

15

10

3

0

1

2

4

5

6

% X 10

RPM

7 8

11

9

10

7

1

0

3

2

4

5

6

~C X 100

8

10

9

8

0

1

4

3

2

5

% X 10

6 7

11

12

10

9

R

% RPM

T

9

7

6

8

45

10000

1000

100

FEE

FEET

FEET

299

2

1

LT TEST

C/W

SELECT CONTROL

VOLTSO.A.T

DAVTRON

TEST

I

SN

T

R

CH

K

D

L

C

E

T

S

T

R

N

H

O

U

E

T

M

FUEL

ON

OFF

TEST

HORNOVSPD

FADEC

FWD TANK

FUEL QTY

VALVE

FADEC

AVOID CONT OPS 68.4% TO 87.1% NP

MODE

LP

FOR

E

R CE

UL

T

QUICK

2 MIN TURN

5

521

10003

2

4

3

1

DOWN

UP

FT

PER MIN

TYPICAL

TEST

FLOAT

70

NR

NP

80

90

100

120

110

60

50

0

10

30

40

20

KNOTS

AIRSPEED

100

120

150

60

80

40

20

0

AUTO

MAN

ON

RESET

E

L

T

A

T

U

O

HDG

6

N

GPS

NAV

21

S

15

12

E

6

3N

33

30

W

24

FLOAT

TEST

FLOAT

ARM BOOST

L/FUELBAGGAGE

DOOR

FADEC

FAULTBOOST

R/FUEL

RELIGHT

AUTO L/FUEL

XFRDOOR

LITTER R/FUEL

XFR FAULT

RESTART

STARTFUEL

FILTEROVERTEMP

HEATER FUEL

VALVE LOW

FUELRPM

OVSPD

OUT

ENGINE

ENGINE

HOT

BATTERY

CYCLIC

CENTERING

HYD

SYSTEM

RLY

TEMP

XMSN OIL

BATTERY

T/R

CHIP

ENGINE

XMSN

CHIP

CHIP

FADEC

MANUAL

FADEC

FADEC

DEGRADED

FAILGEN FAIL

PRESS

INSTR

CHECK

XMSN OIL

1010

30

KING

21

12

3

6

15

24

33

E

S

W

N

GSGS

NAV HDG

ANTI-ICE

ENGINE

21

12

3

33

E

30

15S

24

W

OBS

ENGAGED

PEDAL STOP

PTT

STOP

PEDAL

3

2

4

5

1

8 7 6

407_MD_01_0007

THIS HELICOPTER MUST BE OPERATED INCOMPLIANCE WITH THE OPERATING LIMITATIONS

SPECIFIED IN THE APPROVED FLIGHT MANUAL

FUEL CAPACITYBASIC 869 LBS

WITH AUX 1005 LBS(JET A AT 15° C)

Engine ANTI-ICE segment added

at S/N 53095 and subsequent.

Spare segment on S/N 53000

through 53094.

ELT remote switch

KI-525A horizontal situation

indicator

Airspeed Actuated Pedal Stop

press-to-test switch/annunciator

NAV/GPS switch

KI-208 course deviation indicator

KI-227 ADF indicator

Placard will reflect FADEC

software version installed.

FADEC SOFTWARE VERSION 5.356WITH DIRECT REVERSION TO

MANUAL INSTALLED, REFER TOFLIGHT MANUAL FOR OPERATION.


31 JAN 2007—Rev. 2———1-17

Figure 1-4. Instrument Panel and Pedestal (Sheet 2 of 2)407_MD_01_0008

9.

10.

11.

12.

13.

14.

15.

16.

9

KLN 89B GPS receiver

KMA 24H-71 audio panel

KX-155 or KX-165 VHF

communication/navigation transceiver

KY-196A VHF communication transceiver

KT-70 or KT-76A ATC transponder

KR-87 ADF receiver

KA-51B slave control

ICS mode selector

11

12

13

14

15

16

10


1-18———Rev. 2—31 JAN 2007

Figure 1-5. Overhead Console

HEADPHONE JACKS

TYPICAL

407_MD_01_0009

FWD


31 JAN 2007—Rev. 2———1-19

Figure 1-6. Maintenance Ports407MM_76_0012+

LEFT SIDE LOWER CONSOLE

FADEC/ECU

MAINTENANCE

BUTTON

FADEC/ECU

GROUND

MAINTENANCE

CONNECTOR

INSTALLED WITH

KLN 89B GPS KIT


1-20———Rev. 4—30 APR 2008

1-7. ENGINE INSTRUMENTS

The propulsion instruments are electronicinstruments that are individually poweredthrough circuit breakers located on the 28VDC bus of the helicopter. Each side of dualdisplay instruments such as ENGINE OILPRESSURE and TEMPERATURE are poweredthrough individual circuit breakers.

The dial of the instruments are made up of aseries of single bar LCDs which are arrangedin a trend arc. The last LCD to turn onindicates the reading of the instrument. Thefirst bar indicates power on. Some indicatorsalso include a digital display to provide moreaccuracy of the information displayed.

1-7-A. INSTRUMENT OPERATION

All of the propulsion instruments have abuilt-in-test (BIT) capability.

1-7-A-1. POWER−ON BIT

The power-on BIT starts when the power isapplied to the instrument. The power-on BITdoes an integrity check of the electroniccomponents of the indicator.

NOTE

With the “Ahlers” NR/NP gaugeinstalled, the NR and NP needles donot self test during power-on.

On all indicators except the dual tachometer,during the power-on BIT, the trend arc displayshows the full scale for 6 to 8 seconds.Torque and NG digits display 8188.8. MGT andfuel digits display 81888. NR and NP needlesof the dual tachometer move to 107 and 100%(upper redline limit) respectively.

If an LCD indicator fails the BIT, the trend arcand the digits, if applicable, will show blank. Iffaults are detected in the dual tachometerduring the BIT, pointer movement will notoccur. With the “Ahlers” dual tachometerinstalled, no movement during power-on isnorma l . Fau l ts in the “Ahlers” dua l

tachometer will be displayed by no pointermovement when depressing the LCD TESTbutton. Any failure during the BIT found ontorque, MGT, and NG indicators wil l berecorded in non-volatile memory (NVM) of theinstrument. Instruments that successfullycomplete the BIT will have the first LCD of theindicator lit or will display the appropriateindication.

On the torque MGT and NG indicators, aletter E will appear in the digital display whenthe instrument has recorded an exceedance.Pressing the INSTR CHK button will displaythe value of the last exceedance recorded.Releasing the button will clear the E until thenext power up of the instrument.

1-7-A-2. COMMANDED BIT

The commanded BIT starts when the LCDTEST button is pushed. The commanded BITturns on all the LCDs so that the pilot canverify that all the LCDs are working. On thedual tachometer indicator, the individualRotor and Turbine pointers are driven toindicate their respective upper redline limits.

1-7-A-3. EXCEEDANCE MONITORING

TORQUE, MGT, and NG instruments have anability to record exceedances. Exceedanceshave been predefined at the design stage andpreprogrammed into the microprocessor ineach indicator.

Exceedances a re def ined as l imi ts o foperation above which there may be somemaintenance action required.

NOTE

Exceedance monitoring is providedas an aid to determining requiredmaintenance action. Only the personin command of the helicopter at thetime the exceedance occurs canverify that the exceedance recordedreflects the actual occurrence.

The above instruments have built in NVM thatallows them to store up to 50 exceedance


30 APR 2008—Rev. 4———1-21

events. For each exceedance, the memoryrecords the date, duration, and peak valueduring the exceedance. If exceedance eventsrecorded are greater than 50, the earliestdated exceedance that was in memory isdeleted and the latest exceedance is added inits place.

To provide advance notice to the pilot that anexceedance is about to be recorded, theseindicators also have other preprogrammedadvisory points at which they will flash theTrend ARC display. The digital readoutdisplay will not flash. Specific values arediscussed in applicable systems descriptionsin this section of manufacturer’s data.

When an advisory is disp layed by theinstrument, the instrument will also turn onthe CHECK INSTR light on the caution panel.If the pilot makes control inputs to reduce theinstrument readings below the advisoryvalues, the advisory wi l l no longer bedisplayed and the CHECK INSTR light will beturned off. No exceedance will be recorded.

When the indicator exceeds specificallypreprogrammed values, an exceedance willbe recorded. When an exceedance is about tobe recorded, the CHECK INSTR light will beturned on by the instrument. In addition, whenan exceedance is recorded, the letter E will bedisplayed at the left digit on the digitaldisplay.

The pilot can acknowledge the exceedanceand cause the peak value to be displayed onthe analog and digital display by pushing theINSTR CHK button (Figure 1-4) on theinstrument panel. The exceedance will displayfor a maximum of 11 seconds, less if the pilotpushes on the INSTR CHK button for a shorterper iod o f t ime. I f the re have beenexceedances recorded on different indicators,each ind ica tor w i l l d isp lay i ts las texceedance.

Once the pilot has pushed the INSTR CHKbutton and the exceedance(s) has beendisplayed, the E will disappear from the digitaldisplay. The E will not display until the power

is removed from and applied to the indicator(such as by pulling its circuit breaker orturning helicopter power off and on).

The last exceedance (E) will continue todisplay each time the indicator is powered upuntil the exceedance(s) is removed from theNVM of the indicator using a computer withmaintenance download so ftware(BHT-407-MM).

If an exceedance has not been detected by theTORQUE, MGT, or NG indicator, the CHECKINSTR light and the E on the indicator digitaldisplay will not illuminate when the INSTRCHK button is pressed.

1-8. CHECK INSTRUMENT (CHECKINSTR) LIGHT

The CHECK INSTR light circuit is designed toalert the pilot that either the TORQUE, MGT, orNG indicator is about to or has detected anexceedance.

When the indicators exceed preset values, theindicator's trend ARC display will begin toflash and the CHECK INSTR light will turn on.The digital readout display will not flash. If anE is displayed by an indicator, the CHECKINSTR light will remain on until the pilotacknowledges the exceedance by pushing theINSTR CHK button (Figure 1-4).

1-9. CLOCKThe clock is included in a multifunctionindicator mounted in the upper left area of theinstrument panel. The indicator also displaysOutside Air Temperature (OAT) in °C or °F andvolts (DC).

The clock is a digital display, quartz crystalchronometer.

The clock has a high contrast liquid crystaldisplay and a two button control systemdesigned to prevent accidental time settingwhile selecting various functions. The clockfunctions are as follows:

Universal Time (UT) in 24-hour format.


1-22———Rev. 4—30 APR 2008

Local Time (LT) in 12-hour format(24-hour format may be selected).

Elapsed Time (ET) count up timer tomaximum of 99 hours, 59 minutes.

Elapsed Time (ET) count down timer frommaximum 59 hours, 59 minutes.

Elapsed Time (ET) alarm (flashingdisplay) at zero count down.

Flight Time (FT) count up to a maximumof 99 hours, 59 minutes.

Flight Time (FT) alarm (flashing display)at zero count down.

A line on the display will be positioned overthe letters indicating the function selected(UT/LT/FT or ET).

The clock display receives its power from the28 VDC bus through the OAT/V INSTR circuitbreaker. When the clock is not powered, theCONTROL and SELECT buttons are disabled.The display lighting is powered by 5 VDCinstrument lighting system.

When all electrical power is turned off, theclock display disappears, but the crystaltiming reference continues to operate fromthe power of a 1.5 volt penlight dry cellclipped to the back of the clock case. The drycell is not charged by the helicopter electricalsystems, and it should be replaced annuallyto ensure uninterrupted operation of thetiming reference.

1-9-A. CLOCK OPERATION

The SELECT button selects what is to bedisplayed and the CONTROL button controlswhat is displayed. The SELECT buttonsequentially selects UT, LT, FT, ET and back toUT. The CONTROL button starts and resetsthe elapsed time when momentarily pressed.

The Flight Time is recorded based on 28 VDCpower input received when the helicopterhourmeter is running (paragraph 1-39-B).

1-9-A-1. UNIVERSAL TIME SETTING

Select UT using the SELECT button. Press theSELECT and CONTROL but tonssimultaneously to enter the set mode. Thetens of hours digit will start flashing. TheCONTROL button controls the flashing digitand each push of the button increments thedigit. Once the flashing digit is set, theSELECT button selects the next digit to beset, from left to right across the display. Afterthe last digit is set, a final press of theSELECT button will exit the clock from the setmode. The function indicator will resumenormal flashing to indicate the clock isrunning.

1-9-A-2. LOCAL TIME SETTING

Select LT using the SELECT button. Press theSELECT and CONTROL but tonssimultaneously to enter the set mode. Thetens of hours digit will start flashing. Thesetting operation is the same as UT, exceptthat the minutes are already synchronizedwith UT and cannot be set in LT.

1-9-A-3. ELAPSED TIME COUNT UP

Select ET using the SELECT button. Pressingthe CONTROL button wi l l s tar t the ETcounting up in minutes and seconds untilreaching 59 minutes and 59 seconds. The ETwill then start counting in hours and minutesup to 99 hours and 59 minutes. Pressing theCONTROL button again will reset the ET tozero.

1-9-A-4. ELAPSED TIME COUNT DOWN

Select ET using the SELECT button. Press theSELECT and CONTROL but tonssimultaneously to enter the set mode. Thecount down time can now be set. Entering thetime is identical to UT time setting. When thetime is entered and the last digit is no longerflashing, the clock is ready to start the countdown. Momentarily pressing the CONTROLbutton starts the count down. When the countdown reaches zero, the display will flashand the ET counter will begin counting up.

30 APR 2008 Rev. 4 1-23


Pressing either the SELECT or CONTROLbutton will deactivate the flashing alarm.Pressing the SELECT button will select UTdisplay or pressing the CONTROL button willreset the ET counter to zero.

1-9-A-5. FLIGHT TIME RESET

Select FT using the SELECT button. Press andhold down the CONTROL button for 3 secondsor until the display shows 99:59. Flight timewill be zeroed upon release of the CONTROLbutton.

1-9-A-6. FLIGHT TIME ALARM SET

Select FT using the SELECT button. Press theSELECT and CONTROL but tonssimultaneously to enter the set mode. Thecountdown time can now be set. Entering thetime is identical to UT time setting. When thetime is entered and the last digit is no longerflashing, the clock is ready to start the countdown.

The Flight Time setting has not been affectedby setting the Flight Time Alarm. When theFlight Time begins recording in conjunctionwith the hourmeter (paragraph 1-39-B), theFlight Time alarm will begin its countdown.

1-9-A-7. FLIGHT TIME ALARM DISPLAY

NOTE

When the Flight Time is equal to thealarm time, the display will flash. If FTis not being displayed at the time thealarm becomes active, the clock willautomatically select FT for display.

Press either the SELECT or CONTROL buttonto turn off the alarm. Flight time will remainunchanged and will continue counting.

1-9-A-8. TEST MODE

Hold the SELECT button in for 3 seconds andthe display will read 88:88. This indicates alldigits are functioning properly.

1-10. CAUTION AND WARNINGSYSTEM

The main function of the caution and warningsystem is to detect specified conditions andprovide a visual or visual/audio indication.Detection is accomplished through the use ofmonitoring circuits that, when actuated, causethe annunciator lights to illuminate. Visualindications are provided through a cautionand warning panel while audio indications areprovided through three separate warninghorns: ENGINE OUT (pulsating), LOW ROTOR(continuous), and FADEC FAIL (chime tone)(Figure 1-7).

The caution and warning panel comprises 36individual annunciator posit ions whichprovide a visual indication of cautions (amberlights), warnings (red lights), and advisory(green or white l ights) conditions. Eachannunciator contains three lamps. Individuallamp operation is provided in the event one ormore of the lamps burns out.

Of the 36 annunciators, 32 are illuminated witha ground input (ground seeking), three areilluminated with a 28 VDC bus input (positiveseeking), and one is i l luminated with acombination ground and 28 VDC bus input(ground/positive seeking).

The FLOAT TEST, BATTERY HOT, ENGINEOVSPD, ENGINE OUT, and RPM cannot bedimmed. The remain ing l ights may beilluminated in either the fixed bright or thefixed dim mode. With the instrument lightrheostat knob in the OFF position, all lightswill illuminate in the bright mode if they areturned on. With the instrument light rheostatknob, set between the dim and BRT position,the CAUT LT DIM/BRT switch may be used toselect either the dim or bright mode. In the DIMmode all lights, with the exception of thosementioned above, will be illuminated at a fixeddim value if they are turned on.

1-24 Rev. 2 31 JAN 2007


Figure 1-7. Caution and Warning Panel

1

2

Segment lettering for spare is .

Engine anti-ice segment added at S/N 53095 and subsequent. Spare segment on S/N 53000 through 53094.

407_MD_01_0010

NOTES

1

2


30 APR 2008—Rev. 4———1-25

All of the light lamps can be tested at once bypressing the CAUTION LT TEST switch on theright-hand side of the caution panel. This willalso test the FADEC Mode switch lamps andNAV/GPS, Quiet Mode lamps (if installed). Itwill not test the pedal stop switch lamps. Thepedal stop switch must be pressed (press toTest-PTT) to check its lamps.

1-11. POWER PLANT

The Rolls-Royce 250-C47B engine (Figure 1-8)is a turboshaft engine featuring a free powerturbine. The gas generator is composed of asingle-stage, single entry centrifugal flowcompressor directly coupled to a two-stagegas generator turbine. The power turbine is atwo-stage free turbine which is gas coupled tothe gas generator turbine. The integralreduction gearbox has multiple accessorypads and a splined output shaft which mateswith the freewheel unit. The engine has asingle combustion chamber with single fuelinjection and igniter. The output shaft centerline is located below the center line of theengine and the single exhaust outlet isdirected upward. The engine incorporates aFull Authority Digital Engine Control (FADEC)system. At takeoff power (100% instrumenttorque) , the engine w i l l p rov ide 674horsepower at sea level, 31°C, basic inlet andat maximum continuous power (93.5%instrument torque), the engine will provide630 horsepower at sea level, 19.5°C, basicinlet.

To further clarify the 250-C47B's enginehorsepower, the engine dataplate identifiesthe engine's rated horsepower as 650. Thisrated horsepower is what Rolls-Royceguarantees the engine will provide at sea levelat a specific fuel flow and MGT. This ensuresall 250-C47B engines sold to Bell Helicoptermeet or exceed the rated horsepowerspec i f ica t ion . There fo re , the ra tedhorsepower shown on the engine dataplate isnot the maximum horsepower that the enginewi l l de l iver when insta l led in the 407helicopter.

1-12. ENGINE CONTROLS —FADEC SYSTEM

This paragraph addresses the operationaldesign of the Rolls-Royce 250-C47B FADECsystem, its relationship with airframe androtor systems, possible system faults,troubleshooting, and training procedures. Theinformation provided reflects operation withFADEC software version 5.202 and 5.356 withdirect reversion to manual system installed.A l though F l ight Manua l EmergencyProcedures are explained in this section, referto the BHT-407-FM-1, Section 3, for actualflight operations.

1-12-A. FADEC SYSTEM

The FADEC system is designed to enhanceflight safety and reduce pilot workload as wellas provide other important benefits. Inaddi t ion to the operat iona l benef i t ofincreased TBO, engine automatic start, andprecise control of main rotor speed, the 407fea tures redundant s igna l sens ing ,continuous monitoring, and self diagnostics.Since much of the redundant design istransparent to the pilot, the caution/warning/advisory system has been expanded to adviseof conditions resulting from the increasedmonitoring.

Although the possibility of a FADEC systemfailure is unlikely, pilots and maintenancepersonne l must have an operat iona lunderstanding of the FADEC system, alongwith a sound knowledge of emergency andtroubleshooting procedures. Bell Helicopterrecommends that personnel involved with the407 familiarize themselves with the procedurefor FADEC FAILURE. Familiarization with thisprocedure will help to emphasize FlightManual Emergency Procedures as thefollowing FADEC information is read withinManufacturer’s Data.

A FADEC FAILURE condition during operationin AUTO mode will be indicated by activationof the FADEC FAIL warn ing horn andillumination of the FADEC FAIL and FADECMANUAL warning lights.


1-26———Rev. 2—31 JAN 2007

Figure 1-8. Power Plant407_MD_01_0002

1.

2.

3.

4.

5.

Air inlet

Compressor assembly

Combustion section

Turbine section

Power and accessories gearbox

TYPICAL

3

2

4

5

1


26 MAR 2007—Rev. 3———1-27

Respond to the horn and lights as describedin Sec t ion 3 , Emergency /Ma l func t ionProcedures, in basic Flight Manual.

1-12-B. FADEC SYSTEM — OPERATION

The control system (Figure 1-9) is designed tooperate in many modes, covering all possiblesystem states, including the transitionalmodes to and from the AUTO and MANUALmodes.

The FADEC uses a single channel control withone microprocessor and one electronic lane.There is a lso a MANUAL mode hydromechanical back-up. The FADEC system hastwo main components: the airframe mountedElectronic Control Unit (ECU) and the enginemounted Hydro mechanical Unit (HMU).

The ECU monitors numerous internal andexternal inputs to modulate fuel flow andtherefore control engine speed, accelerationra te , temperatu re , and other eng ineparameters. The ECU provides inputs to theHMU to modulate fuel flow based on thecontinuous monitoring of the following:Measured Gas Temperature (MGT), GasProducer speed (NG), Power Turbine speed(NP), Main Rotor speed (NR), Engine TorqueMeter Oil Pressure (TMOP), Collective Pitch(CP) and rate, Compressor Inlet Temperature(CIT), Ambient Pressure (P1), and PowerLever Angle (PLA)/throttle position.

In addition, when 5.356 software is installed,the engine uses a digital electronic controlsystem based on two electronic governorscalled primary channel and reversionarygovernor. The p r imary channe l is afull-authority digital electronic control(FADEC) that controls, monitors, and limitsengine power while maintaining helicopterrotor speed. The reversionary governor canautomatically take control of the engine in theevent of a primary channel failure. Thereversionary governor uses a limited set ofinputs and prov ides bas ic e lect ronicgoverning.

The HMU consists of a two stage suction fuelpump, fuel metering assembly, Auto/Manualchangeover so lenoid va lve , e lect r icoverspeed valve, mechanical fuel shut offvalve, hot start fuel solenoid valve, andaltitude compensated bellows. The HMUprovides fuel modulation via a stepper motorin AUTO mode and a hydro mechanicalactuator in MANUAL mode.

1-12-C. POWER UP MODE ANDBUILT-IN-TEST

The FADEC system incorporates logic andcircuitry to perform self-diagnostics. Ingeneral, sensors are checked for continuity,rate and proper range. Discrete inputs arechecked for continuity and output drivers aremonitored for current demand to sense failedactuators and open or shorted circuits. AFADEC power up check exercises outputdrivers and actuators to ensure systemfunctional i ty and readiness. The briefappearances of light indications and theirrespective horn observed immediately afterapplication of power are normal and are partof the FADEC system’s designed initializationprocess. If any faults are detected during theself-test, the appropriate FADEC cautionpanel light will illuminate.

The helicopter 28 VDC bus supplies electricalpower to the FADEC ECU until the engineachieves 85% NP. Above this speed, theFADEC ECU will select between the 28 VDCbus and the engine-driven permanent magnetalternator (PMA), as its primary power source.The higher voltage source will be selected. Inthe event of a primary power source failure,the alternate source will be selected.

1-12-D. START IN AUTO MODE

To ready the system for an automatic start,the FADEC MODE switch must be set toAUTO, and the throttle set to the idle position.The s tar t sw i tch is then momentar i lypositioned to START. Observe START andAUTORELIGHT lights are illuminated beforereleasing START switch. Throttle modulationof fuel flow is not required.


1-28———Rev. 3—26 MAR 2007

Figure 1-9. FADEC Control System Schematic407MM_76_0001_c01+

OIL PRESSURETORQUE METER

(TMOP)SPEED (N )POWER TURBINE

SPEED (N )

GAS GENERATOR

COMPRESSOR

TEMPERATURE

(CIT)

PERMANENT MAGNET

N

N

HYDRO-MECHANICAL

FUEL NOZZLE

FUEL IN

LEVERPOWER

ENGINE

AIRFRAME

THROTTLE(PLA) LINKAGE

FUEL

UNIT (ECU)

ELECTRONIC

AMBIENTPRESSURE(P1)

ISOLATORSVIBRATION

AUTO RELIGHTIGNITER RELAY

STARTER RELAYLATCHING

MAIN ROTOR

+28 VDC BATTERY/

FADEC/ECU MTCE

MTCE TERMINAL

COCKPIT TYPICAL

N

N /N

MGT

OVERSPEED

ON

MANUAL

PITCH (CP)

AUTO

FUELFILTER

OUT

UNIT(HMU)

THROTTLEPOWER LEVERANGLE (PLA)

-FADEC FAIL-FADEC MANUAL-FADEC DEGRADED-FADEC FAULT-RESTART FAULT-ENGINE OVSPD-ENGINE OUT-AUTO RELIGHT

CAUTION

PANEL LIGHTS

AND RATE

COLLECTIVE

FADEC MODE SWITCH

TEST SWITCH

120

60

60

120

(MECHANICAL LINKAGE)

CONNECTOR)

PORT (EMC-35A

AIRFRAME POWER

SPEED (N )

CONTROL

(PMA)

ALTERNATOR

INLET(MGT)

MEASURED

GAS

TEMPERATURE

G

P R

R

P

G

G

P


31 JAN 2007—Rev. 2———1-29

Although the start sequence is automatic, thepilot is responsible for monitoring the startand taking appropriate action if required.Therefore, it is recommended that both thethrottle and start switch are guarded until thestart is completed. Do not initiate a start ifFADEC related caution panel l ights areilluminated unless appropriate maintenanceinvestigation or successful corrective actionhas been carried out and no "current" faultsare shown (paragraph 1-12-R-2).

NOTE

After the throttle is set to idle, themomentary contact start switch mustbe activated within 60 seconds toinitiate the start and engage thelatching feature. The latching featureof the start will engage when theFADEC ECU senses momentaryactivation (1 second) of the startswitch or upon sensing an NG speedof 5%. If a start is attempted followinga delay of more than 60 seconds, theFADEC system will not allow thestarter to latch following release ofthe s tar t sw i tch , and w i l l notintroduce fuel if the start switch isheld to START. Therefore, if a delayof more than 60 seconds hasoccurred, the system must be reset.To reset the system, the throttle mustbe repositioned to cutoff and thenback to idle. In addition, if electricalpower is in ter rupted pr io r toinitiating the start, with the throttle atid le , the thro t t le must berepositioned to cutoff and then backto idle after power is restored tore-enable the latching feature. Anormal automatic start sequencemay then commence. If startingengine on external power, refer toparagraph 1-32-A, External Power.

To allow for cooler starts and reduce thepossibility of reaching hot start abort limits, it

is recommended that residual MGT be below150°C, when below 10,000 feet HP or below65°C, when above 10,000 feet HP prior to start.To reduce residual MGT, a Dry Motoring Runmay be performed in accordance with theFlight Manual.

Activating the automatic start mode engagesthe airframe mounted FADEC/start relaywhich is then latched by the FADEC ECU untilthe NG speed reaches 50%. The FADEC/startrelay places the MGT indicator into the startmode, signals the generator control unit/voltage regulator to inhibit generator output,flashes the shunt field for the duration of thestart, and activates the starter relay. Thestarter relay activates the starter, illuminatesthe START advisory segment on the cautionpanel, and activates the igniter relay. Theigniter relay activates the engine ignitersystem and illuminates the AUTO RELIGHTadvisory segment of the caution panel. Whilein s tar t mode, the MGT ind ica tor isprogrammed to record s ta r t MGTexceedances, should they occur, which differfrom normal operational MGT exceedancelimits.

NOTE

The fo l lowing paragraph isapplicable if helicopter is configuredwith FADEC Software Version 5.356(Reversionary Governor).

Once NG speed reaches 10% for ambienttemperatures of 26.6°C (80°F) or below, or12% for ambient temperatures above 26.6°C(80°F), the FADEC system will introduce fuel,detect the lightoff, and smoothly acceleratethe engine to idle while limiting MGT ifnecessary. At inlet temperatures below -18°C,the FADEC will increase start accelerationfrom 2 to 6% NG per second. The increase inthe acceleration rate will be noticeable duringstart and improves start performance atcolder temperatures and at higher altitudes.


1-30———Rev. 2—31 JAN 2007

NOTE

The fo l lowing paragraph isapplicable if helicopter is configuredwith FADEC Software Version 5.202.

Once NG speed reaches 10% for ambienttemperatures of -6.7°C (20°F) or below, or 12%for ambient temperatures above -6.7°C (20°F),the FADEC system will introduce fuel, detectthe lightoff, and smoothly accelerate theengine to idle while limiting MGT if necessary.At inlet temperatures below -18°C, the FADECwill increase start acceleration from 2 to 6%NG per second . The increase in theacceleration rate will be noticeable duringstart and improves start performance atcolder temperatures and at higher altitudes.

Upon reaching an engine NG speed of 50%,the FADEC ECU unlatches the FADEC/startrelay, terminating the start sequence. Inaddi t ion , the ECU increments thestart-counter to the next number when lightoffis detected.

Above 50% NG, the FADEC ECU carries out aself test of the auto relight system andcontinues to energize the igniter relay until 60±1% NG. During this time, the AUTO RELIGHTlight will be illuminated. The engine willcontinue to accelerate until reaching astabilized idle of 63 ±1% NG.

The FADEC system also incorporates "HotStart Abort Logic" up to 50% NG during start.This ECU controlled feature will cut off fuelflow to the engine fuel nozzle if any of thefollowing conditions occur:

Start MGT exceeds 843°C, at pressurealtitudes less than 10,000 feet and if ECUdetermined residual MGT was less than82.2°C at initiation (5% NG) of start.

Start MGT exceeds 912°C, at pressurealtitudes greater than 10,000 feet or ifECU determined residual MGT wasgreater than 82.2°C at initiation (5% NG)of start.

Voltage to FADEC ECU drops below 10.3VDC. As a significant momentary voltagedrop occurs at initiation of the start,ensuring a battery voltage of 24 VDC orabove prior to start, in conjunction withappropriate battery maintenance, willreduce the poss ib i l i ty o f vo l tagedropping to 10.3 VDC.

If a FADEC aborted start occurs or the pilotmanually initiates a start abort by positioningthe throttle to cut off, the FADEC is designedto automatically keep the starter engaged forup to 60 seconds from initiation of the start, toreduce MGT to 150°C. Once 150°C MGT isobtained, the starter will disengage.

NOTE

Momentarily positioning the startswi tch to D ISENG wi l l on lydeactivate the FADEC/start relaywhich in turn disables the starter andign i te r c i rcu i ts . In the eventdeactivation of the starter and ignitercircuits occurs after engine light-off,but below 50% NG, the FADEC ECUwill either modulate fuel flow toprov ide a star t i f N G speed issufficient, or cutoff fuel flow if MGTexceeds hot start abort limits due tolow NG speed. Therefore, positioningthe th rot t le to cuto f f is theappropriate method to manually stopthe start sequence.

Following start and prior to increasing enginespeed and main rotor RPM to 100%, a FADECMANUAL check is required. The purpose ofthis check is to ensure the engine responds tothrottle movement while in MANUAL mode.Prior to positioning the FADEC mode switchto MANUAL for the purposes of this check,ensure the throttle is positioned to idle andNG speed is at 63 ±1%.

NOTE

The engine idle speed may reduce tothe point where the engine out lightand horn are activated.


31 JAN 2007—Rev. 2———1-31

NOTE

If NG increases to more than 75% NGwith throttle positioned in idle detent,maintenance action is required.Refer to the Rolls-Royce 250-C47BOperation and Maintenance Manual.

Once the FADEC mode switch is positioned toMANUAL, the FADEC MANUAL and AUTORELIGHT lights will illuminate. In addition, theNG speed may change from 63 ±1%. Anincrease or decrease in NG speed may benoticed. The change in NG speed is due to thefact that the FADEC system is nottemperature compensated in regard to fuelflow in MANUAL mode. Engine load and HMUcalibration will also play a part in determiningif the NG speed will change.

Once NG is stabilized, slowly increase throttleto approximately 5 to 10% NG to ensureengine responds and then return throttle toidle position. Position FADEC Mode switch toAUTO and ensure FADEC MANUAL and AUTORELIGHT lights extinguish. Engine shouldreturn to an NG speed of 63 ±1%.

Engine acceleration from idle to 100% NR/NP,in AUTO mode, is achieved by smoothlyincreasing the throttle to the FLY detentpos i t ion (PLA 70° ) . As the thro t t le ispositioned from idle (PLA 30° to 40°) to theFLY detent position (PLA 70°), electricalsignals are sent to the ECU from the HMU –PLA potentiometer. These signals dictate theamount of authority the ECU has to controlmaximum fuel flow (NG limiting) based onthrottle position, and in turn controls engineNG speed. Therefore, as the throttle isincreased from idle to the FLY detent position,the fuel flow is electronically increased until100% NR/NP is obtained. To avoid rapid engineacceleration, it is recommended that throttleapplication from idle to the FLY detentposition be conducted in a smooth andgradual manner.

1-12-E. ALTERNATE START — AUTOMODE

For helicopters that operate at temperaturesof approximately 26.6°C (80°F) and above, orat high altitudes and experience a hot startabort event, the start procedure listed belowis approved and has been demonstrated toovercome this issue in most instances.

For hot and/or high altitude environments,and when prior troubleshooting has notrevealed any engine maintenance issues, thisprocedure can be used to alleviate thepossibility of aborted hot starts.

With the collective in the full down positionand the throttle in CUTOFF, the pilot caninitiate the start by pressing the starterswitch. This will energize the starter motorand turn on the ignition exciter (continue tohold starter switch). Once N1 has reachedapproximately 16%, the pilot is to move thethrottle from CUTOFF to IDLE. Once throttle ismoved to IDLE position, the starter will latchand the starter switch can be released oncelight off is detected. The engine will detectlight off and smoothly accelerate to groundidle while limiting the MGT if necessary.

1-12-F. START IN MANUAL MODE

In accordance with the Rolls-Royce 250-C47BOperation and Maintenance Manual, ManualMode starting on the ground is not authorizedexcept for use under emergency conditions orunder special permit from the local aviationauthority. Refer to the Rolls-Royce 250-C47BOperation and Maintenance Manual to ensureall Manual Mode Operational Procedures arefollowed. Automatic hot start abort featuresare not available in MANUAL mode.

To ready the system for a manual start, theFADEC MODE switch is set to MAN and thethrottle positioned to cutoff. In MANUAL, theFADEC will not latch the FADEC/start relay orcontrol the fuel scheduling during the start.The start switch must be held in the STARTposition until the start sequence is completed


1-32———Rev. 2—31 JAN 2007

at 50% NG. When the NG speed reaches 12%,or NG speed speci f ied in Rol ls-Royce250-C47B Operation and Maintenance Manualfor specific OAT, slowly advance the throttleout of cutoff and stop when the engine lightsoff. Allow the MGT to peak and then increasefuel flow by modulating throttle to maintainMGT within limits.

Once the engine has been started, positionthe throttle to the idle detent and monitor theNG speed.

NOTE

The engine idle speed may reduce tothe point where the engine out lightand horn are activated.

NOTE

If NG increases to more than 75% NGwith throttle positioned in idle detent,maintenance action is required.Refer to the Rolls-Royce 250-C47BOperation and Maintenance Manual.

Id le detent speed in MANUAL may notstabilize at 63 ±1% NG.

• If NG speed stabilizes below 63 ±1%,adjust throttle to maintain 63 ±1% NG.

• If NG speed stabilizes between 63 ±1%and 75% this is acceptable provided NPspeed does not fall into avoid STEADYSTATE range of 68.4 to 87.1% NP. If thisoccurs, adjust throttle to avoid 68.4 to87.1% NP.

• If NG stabilizes above 75% NG,maintenance action is required.

Once a MANUAL mode star t has beeninitiated, it may be terminated at any time byrotating the throttle to cutoff. The pilot shouldcontinue motoring the engine until the MGThas stab i l ized a t an acceptable leve l .Releasing the start switch from the STARTposition will disengage the FADEC/start relayand disable the starter and igniter circuits.

After a successful MANUAL mode start, andidle has been achieved and is stable, performa momentary switch back to AUTO mode. Ifengine flameout occurs, subsequent starts/flight are prohibited until appropriate FADECsystem troubleshooting has been performed.

Engine acceleration from idle to 100% isachieved by positioning the throttle towardsfu l l open. Th is is ach ieved th roughhydromechanical control of the HMU fuelmetering valve.

1-12-G. IN-FLIGHT AUTO MODEOPERATION

NOTE

If throttle is not maintained in FLYdetent position during normal flightoperations, available engine powermay be limited. This will occur ifthrottle is positioned from the FLYdetent (PLA 70°) to a setting which isless than 62° PLA. Although anythrottle position from 62° PLA to fullopen will provide the FADEC withcomplete control to maintain NRwithin limits, the approved throttleposition for flight is the FLY detentposition (PLA 70°).

During flight in AUTO mode with the throttlein FLY detent position (PLA 70°), the FADEChas complete control over engine operation tomaintain NR within limits. The ECU receivesengine and airframe parameter inputs andcockpit command signals, processes them,and modulates the HMU stepper motor drivenfuel metering valve to achieve desired engineperformance.

If required, as may be the case in certainEmergency Procedures, an alternate meansof engine control is also available to the pilotin AUTO mode. This can be achieved bymanipulating the throttle below FLY detentpos i t ion unt i l the requ i red eng ineperformance is achieved. As the throttle ispositioned between HMU Power Lever Angles


31 JAN 2007—Rev. 2———1-33

of 40 to 62°, electrical signals are sent to theECU from the HMU-PLA potentiometer. Thesesignals dictate the amount of authority theECU has to control maximum fuel flow (NGlimiting), and in turn, engine NG speed.Therefore, as throt t le is increased ordecreased, the maximum NG speed isregulated electrically by limiting the fuel flow.

In AUTO mode, the FADEC is capable ofdetecting an engine flameout by sensing NGdeceleration. Without any pilot action, theau to - re l igh t sequence is in i t ia ted byestablishing a control led fuel f low andactivating the ignition system. The FADEC willcontrol the MGT and accelerate the engine tothe previously selected state.

The automatic relight sequence will initiate atdetection of a flameout, continuing theprocedure until a relight occurs or the NGspeed decays to 50%. During this time, theAUTO RELIGHT and ENGINE OUT cautionpanel lights will be illuminated and the engineout warning horn will be activated. If the NGdecays below 50%, the FADEC system willdiscontinue the attempt to relight the engineand the ENGINE OUT light and warning hornwi l l remain act ive . In the event o f anunsuccessful relight, refer to the restart –automatic mode emergency proceduredescribed in the BHT-407-FM-1. For additionalinformation on in-flight AUTO mode restarts,refer to paragraph 1-18.

While in AUTO mode, if any failure occurs inthe ECU or in one of its input/output signalsthat significantly impacts the ECU control ofthe HMU, the pilot will be alerted via theFADEC FAIL warning horn and the FADECFAIL and FADEC MANUAL warning lights. Ifthe detected failure does not significantlyimpair the functioning of the ECU, the pilotwill be alerted via a FADEC DEGRADED,FADEC FAULT caution light, RESTART FAULTadvisory light, or combination of, dependingon the nature of the fault.

1-12-H. FADEC SYSTEM FAULTS

CAUTION

BELL HELICOPTER REQUIRESMAINTENANCE ACTION, PRIOR TOFLIGHT, WHEN A FADEC RELATEDLIGHT IS ILLUMINATED.

There are eight lights in the caution/warning/advisory panel that are controlled by theFADEC: FADEC FAIL, FADEC MANUAL,FADEC DEGRADED, FADEC FAULT, RESTARTFAULT, ENGINE OVSPD, ENGINE OUT, andAUTO RELIGHT.

The FADEC ECU continuously monitors theFADEC system for fau l ts and makesappropriate accommodations to continueoperat ion . Fau l t codes have beenpreassigned to those parameters beingmonitored by the FADEC ECU.

Faults and exceedances can be recordedunder the following conditions:

Engine operating:

When the engine is operating (i.e.,Iightoff has been detected) theFADEC will automatically recordfaults/exceedances as current, lastengine run, and accumulated, asthey occur.

Engine not operating:

When the engine is not operating, butelectrical power is applied, theFADEC will only record current faultsas they occur.

If any failure occurs in the ECU/HMU or in oneof the input/output signals that significantlyimpacts the ECU or control of the HMU, thepilot will be alerted via the FADEC FAILwarning horn and the FADEC FAIL/FADECMANUAL warning lights.


1-34———Rev. 2—31 JAN 2007

With FADEC Software Version 5.356 installed,the reversionary (backup) governor will beactivated under certain fault conditions toeliminate a FADEC FAIL condition. This willallow operations in a degraded mode whileremaining in AUTO mode (paragraph 1-12-M).

If the detected failure does not significantlyimpair the functioning of the ECU, the pilotwill be alerted via a FADEC DEGRADED,FADEC FAULT caution light, RESTART FAULTadvisory light, or a combination of, dependingupon the nature of the fault.

If the fault is minor in nature, it will not becommunicated to the pilot with the enginerunning. These faults are identif ied asmaintenance advisory faults and will bedisplayed during shutdown when the throttleis placed in the cutoff position and NG speeddecays below 9.5%. This will be in the form ofa FADEC DEGRADED light.

The BHT-407-FM-1 provides the appropriateaction required by the pilot for each light orlight/horn condition.

All FADEC faults have been categorized intofive types. The first four relate to in-flightfaults and the fifth relates to MaintenanceAdvisory faults with the engine shut down.Maintenance Advisory faults displayed duringshutdown wil l be discussed under theheading FADEC SYSTEM FAULTS — ENGINESHUT DOWN.

1-12-I. CATEGORY 1 — FADEC FAIL/FADEC MANUAL — DIRECTREVERSION TO MANUALSYSTEM (DRTM)

Faul ts that requ i re p i lo t ac t ion andautomatically initiate a transition to theMANUAL mode will be displayed immediatelywhen detected by the ECU. These faults willdisplay as FADEC FAIL/FADEC MANUAL. TheFADEC FAIL horn (chime tone) will activate inconjunction with the FADEC FAIL and FADECMANUAL warning lights. In addition, theRESTART FAULT light will also be displayed

with a FADEC FAIL /FADEC MANUALcondition.

The DIRECT REVERSION to MANUAL systemensures all FADEC failures revert directly toMANUAL. FAIL FIXED failures do not existwithin this system. In addition, the systemincorporates a throttle which is detented atthe 90% NG bezel FLY position.

The main intent of DIRECT to MANUALsystem is to simplify pilot procedures in theevent o f a FADEC fa i lure . Th is isaccomplished by allowing the pilot to keephis hands on the controls during a FADECfailure and enable an increase or decrease inthrottle from the FLY detent position, asrequired. The pilot will only have to removehis hand from the collective to press theFADEC MODE switch to silence the horn onceestablished in MANUAL mode.

1-12-I-1. THE FADEC SYSTEM WILL FAILTO THE MANUAL MODE ASFOLLOWS:

In this situation, reversion to MANUAL modewill occur, independent of the position of theFADEC MODE switch on the instrument panel.The reversion to MANUAL mode will beginimmediately.

This condition can be identified by thesounding of the FADEC FAIL warning horn,illumination of the FADEC FAIL and FADECMANUAL caution panel lights and FADECMODE switch MAN light at the time of thefailure. As the FADEC SYSTEM has initiatedthe transition to MANUAL mode, the pilotmust be aware that an increase or decrease inNR/NP may occur within 2 to 7 secondsfollowing a direct failure to MANUAL. If thisoccurs, collective and throttle will have to beused to control RPM.

The objective of the DIRECT REVERSION toMANUAL system is to provide the pilot withthrottle control of the fuel metering valve in atimely manner. MANUAL mode allows the


31 JAN 2007—Rev. 2———1-35

pilot to control NR/NP with coordinatedcontrol of the collective and throttle.

The following procedural steps are required:

1. Throttle — If time permits, matchthrottle bezel position to NGindication.

This has proven to reduce the possibilityof an overspeed o r underspeedcondition.

This procedure also permits a smoothertransition to MANUAL mode. This is dueto the fact that the actual NG speed andfuel metering valve position prior toswitching to MANUAL mode will be veryclose to that following the transition toMANUAL mode, resulting in little, if any,RPM change.

If time does not permit matching ofthrottle bezel position to NG indication,initially maintain throttle in detented FLYposition.

2. NR/NP — Maintain 95 to 100% withcollective and throttle.

It is most important to ensure that NR/NP ismonitored and properly controlled, duringand following the transition to MANUALmode.

Within 2 to 7 seconds after FADEC FAILwarning, NR/NP may increase or decreasevery rapidly. This will require collectiveinputs to control RPM. The 407 rotorsystem is very responsive to collectiveinputs and can be controlled by the pilotshould a NR/NP overspeed/underspeedtendency arise.

As it takes 2 to 7 seconds to complete thetransition to MANUAL mode, use of

thro t t le to cont ro l NR /N P w i l l beineffective until the transition to MANUALmode is complete. The transition will notbe completed until the fuel meteringvalve in the HMU can be manual lycontrolled by the pilot through use of thethrottle on the collective.

There are two pistons within the HMU(Figure 1-10): a Manual Load Piston (slowpiston) and a PLA Follower Piston (fastpiston), which must hydromechanicallyextend to contact opposite sides of thefuel metering valve shaft lever. The twopistons move at different rates towardthe fuel metering valve lever. It takesapproximately 2.0 seconds for bothpistons to make contact with the fuelmeter ing va lve lever fo l lowing atransition from an initial condition of lowfuel flow. Similarly, up to 7 seconds maybe required for the two pistons to makecontact following a transition from aninitial condition of high fuel flow. Refer toFigure 1-10 through Figure 1-12 foraddit ional information on AUTO toMANUAL mode transitions.

An increase in NR/NP speed may beexper ienced while in t ransi t ion toMANUAL from a condition of low tohigher fuel flow or high fuel flow to ahigher fuel flow. This will be seen if thethrottle to bezel selection made by thepilot in step 1 of the procedure is higherthan the actual NG speed at the time ofthe FADEC FAILURE condition.

This will occur during the period whenthe HMU Manual Load Piston (slowpiston) engages the fuel metering valvelever and moves i t to a more openposition until the PLA Follower Piston(fast piston) is contacted.


1-36———Rev. 2—31 JAN 2007

Figure 1-10. Auto to Manual Transition at Low Fuel Flow (Sheet 1 of 2)407MM_76_0006+

HIGHPRESSURE

FUEL

2.0 SECONDS

1.0 SECOND

METERINGVALVELEVER

MIN FUELFLOW

MAX FUELFLOW

VARIABLEORIFICE

PLA FOLLOWERPISTON (FAST)

MANUAL LOADPISTON (SLOW)

AUTO/MANUALCHANGEOVER

SOLENOID VALVE

VALVESHOWNOPEN

(DE-ENERGIZED)

POWERLEVER

LOWPRESSURE

FUEL

PARTIAL CROSS SECTION OF HMU

2

1


31 JAN 2007—Rev. 2———1-37

Figure 1-10. Auto to Manual Transition at Low Fuel Flow (Sheet 2 of 2)407MM_76_0007+

2

AUTO TO MANUAL TRANSITION AT LOW FUEL FLOW

INITIAL CONDITION:

Auto mode, low engine fuel flow, throttle at "FLY'' position.

Auto/manual changeover solenoid valve is de-energized, allowing high pressure fuel to manual

load piston and PLA follower piston.

Manual load piston "slowly'' extends. Engages metering valve lever in approximately 2.0 seconds

and begins to drive it to meet PLA follower piston.

Concurrently, PLA follower piston "rapidly'' extends in approximately 1.0 second to a position

that is a function of throttle (PLA) position.

Matching throttle and bezel to the actual N speed will allow the PLA follower piston to position

itself very close to the actual position of the metering valve at the time of the transition. This will

minimize the fuel flow change during the transition. Positioning the throttle to a bezel setting that

is higher than actual N speed at the time of the transition will produce an increase in fuel flow

during the transition. The increase in fuel flow will be caused as the manual load piston engages

the metering valve lever and drives it towards the PLA follower piston. Inversely, positioning the

throttle to a bezel setting that is lower than actual N speed at the time of the transition will

produce a decrease in fuel flow during the transition. The decrease in fuel flow will be caused as

the PLA follower piston engages the metering valve lever and drives it towards the manual load

piston.

After both pistons engage, manual mode is established and no delay exists between throttle

(PLA) movement and fuel flow change.

Slew rate limiting is achieved by hydraulic dynamics.

After FADEC mode switch manual selection or initiation of direct reversion to manual:

Auto/manual changeover solenoid valve normally closed (energized) in auto mode.

Auto/manual changeover solenoid valve is opened (de-energized) for transition to and

during manual mode operation. With the valve open, fuel pressure is used to position the

manual load piston and PLA follower piston.

PLA follower piston is controlled by throttle (PLA) position during transition to manual

and when in manual mode. PLA follower piston position is regulated by fuel pressure

bleed through variable orifice. This provides the means to increase or decrease fuel flow

by altering the position of the fuel metering valve. Manual load piston ensures metering

valve lever is held against PLA follower piston.

When in automatic mode, both the PLA follower piston and manual load piston are

retracted from the metering valve lever. They are held in the retracted position by fuel

pressure when the auto/manual changeover solenoid valve is closed (energized).

NOTES

G

G

G

1


1-38———Rev. 2—31 JAN 2007

Figure 1-11. Auto to Manual Transition at Intermediate Fuel Flow (Sheet 1 of 2)407MM_76_0008+

METERINGVALVELEVER

MIN FUELFLOW

MAX FUELFLOW

0.5 SECOND

3 SECONDS

HIGHPRESSURE

FUEL

VARIABLEORIFICE




SOLENOID VALVE

VALVESHOWNOPEN

(DE-ENERGIZED)

POWERLEVER

LOWPRESSURE

FUEL


2

1


31 JAN 2007—Rev. 2———1-39

Figure 1-11. Auto to Manual Transition at Intermediate Fuel Flow (Sheet 2 of 2)407MM_76_0009+

AUTO TO MANUAL TRANSITION AT INTERMEDIATE (CRUISE) FUEL FLOW

INITIAL CONDITION:

Auto mode, intermediate engine fuel flow, throttle at "FLY'' position.

Auto/manual changeover solenoid valve is de-energized allowing high pressure fuel to manual

load piston and PLA follower piston.

Manual load piston "slowly'' extends. Engages metering valve lever in approximately 3.0 seconds

and begins to drive it to meet PLA follower piston.

Concurrently, PLA follower piston "rapidly'' extends in approximately 0.5 second to a position

that is a function of throttle (PLA) position.

Matching throttle and bezel to the actual N speed will allow the PLA follower piston to position

itself very close to the actual position of the metering valve at the time of the transition. This will

minimize the fuel flow change during the transition. Positioning the throttle to a bezel setting that

is higher than actual N speed at the time of the transition will produce an increase in fuel flow

during the transition. The increase in fuel flow will be caused as the manual load piston engages

the metering valve lever and drives it towards the PLA follower piston. Inversely, positioning the

throttle to a bezel setting that is lower than actual N speed at the time of the transition will

produce a decrease in fuel flow during the transition. The decrease in fuel flow will be caused as

the PLA follower piston engages the metering valve lever and drives it towards the manual load

piston.

After both pistons engage, manual mode is established and no delay exists between throttle

(PLA) movement and fuel flow change.




Auto/manual changeover solenoid valve is opened (de-energized) for transition to and

during manual mode operation. With the valve open, fuel pressure is used to position the

manual load piston and PLA follower piston.

PLA follower piston is controlled by throttle (PLA) position during transition to manual

and when in manual mode. PLA follower piston position is regulated by fuel pressure

bleed through variable orifice. This provides the means to increase or decrease fuel flow

by altering the position of the fuel metering valve. Manual load piston ensures metering

valve lever is held against PLA follower piston.

When in automatic mode, both the PLA follower piston and manual load piston are

retracted from the metering valve lever. They are held in the retracted position by fuel

pressure when the auto/manual changeover solenoid valve is closed (energized).

NOTES

G

G

G

1

2


1-40———Rev. 2—31 JAN 2007

Figure 1-12. Auto to Manual Transition at High Fuel Flow (Sheet 1 of 2)407MM_76_0010+

METERINGVALVELEVER

MIN FUELFLOW

MAX FUELFLOW

0.1 SECOND

6 TO 7 SECONDS

HIGHPRESSURE

FUEL

VARIABLEORIFICE




SOLENOID VALVE

VALVESHOWNOPEN

(DE-ENERGIZED)

POWERLEVER

LOWPRESSURE

FUEL


2

1


31 JAN 2007—Rev. 2———1-41

Figure 1-12. Auto to Manual Transition at High Fuel Flow (Sheet 2 of 2)407MM_76_0011+

AUTO TO MANUAL TRANSITION AT HIGH FUEL FLOW

INITIAL CONDITION:

Auto mode, high engine fuel flow, throttle at "FLY'' position.

Auto/manual changeover solenoid valve is de-energized allowing high pressure fuel to manual load piston

and PLA follower piston.

Manual load piston "slowly'' extends. Engages metering valve lever in approximately 6 to 7 seconds and

begins to drive it to meet PLA follower piston.

Concurrently, PLA follower piston "rapidly'' extends in approximately 0.1 second to a position that is a

function of throttle (PLA) position.

Matching throttle and bezel to the actual N speed will allow the PLA follower piston to position itself very

close to the actual position of the metering valve at the time of the transition. This will minimize the fuel

flow change during the transition. Positioning the throttle to a bezel setting that is higher than actual N

speed at the time of the transition will produce an increase in fuel flow during the transition. The increase

in fuel flow will be caused as the manual load piston engages the metering valve lever and drives it

towards the PLA follower piston. Inversely, positioning the throttle to a bezel setting that is lower than

actual N speed at the time of the transition will produce a decrease in fuel flow during the transition. The

decrease in fuel flow will be caused as the PLA follower piston engages the metering valve lever and

drives it towards the manual load piston.

After both pistons engage, manual mode is established and no delay exists between throttle (PLA)

movement and fuel flow change.




Auto/manual changeover solenoid valve is opened (de-energized) for transition to and during manual

mode operation. With the valve open, fuel pressure is used to position the manual load piston and PLA

follower piston.

PLA follower piston is controlled by throttle (PLA) position during transition to manual and when in

manual mode. PLA follower piston position is regulated by fuel pressure bleed through variable orifice.

This provides the means to increase or decrease fuel flow by altering the position of the fuel metering

valve. Manual load piston ensures metering valve lever is held against PLA follower piston.

When in automatic mode, both the PLA follower piston and manual load piston are retracted from the

metering valve lever. They are held in the retracted position by fuel pressure when the auto/manual

changeover solenoid valve is closed (energized).

NOTES

1

2

G

G

G


1-42———Rev. 2—31 JAN 2007

Inversely, a decrease in NR/NP speed maybe experienced during the transition toMANUAL from a condition of low to lowerfuel flow or high fuel flow to a lower fuelflow. This will be seen if the throttle tobezel selection made by the pilot instep 1 of the procedure is lower than theactual NG speed at the time of the FADECFAILURE condition. This will occurduring the period when the PLA FollowerPiston (fast piston) engages the fuelmetering valve lever and moves it to amore closed position as dictated bythrottle to bezel position.

The approximate time to detect a powerchange during the transition to manual issummarized in Table 1-1.

In simpler terms, there will be a timedelay and possible change in enginepower while the system transitions toMANUAL. The length of the delay anddegree of power change during thetransition depends on engine power atthe time of the FADEC failure and the

desired power as selected by throttleposition for the transition. As statedpreviously, the degree of power changecan be minimized by matching thethrottle bezel to the actual indicated NGspeed in step 1 of the FADEC FAILUREprocedure. This permits the smoothertransition to MANUAL mode due to thefact that the actual NG speed and fuelmeter ing va lve pos i t ion p r ior toswitching to MANUAL will be very closeto that fo l lowing the t ransi t ion toMANUAL mode. This will result in little, ifany, power/RPM change.

Once both pistons contact the lever onthe fue l meter ing va lve shaft , thetransition to MANUAL Mode will becomplete. The pilot will have slew ratelimited control of the fuel metering valvevia throttle position without any delay.

Throttle may now be used, in conjunction withcollective, to maintain rotor and engine RPMwithin 95 to 100%.

Once in MANUAL mode, the pilot will havecomplete control of NR/NP by flight controlmanipula t ion and the thro t t le on theco l lect ive . The fue l f low s lew ra te ishydromechanically limited to provide properresponsiveness for helicopter operation andalso prevents surge. Fuel flow will be afunction of the pilot controlled fuel meteringvalve orifice size. Maximum ContinuousPower wil l be avai lable for al l ambient

conditions. MANUAL mode fuel f low ispressure altitude compensated to maintain anapproximate constant horsepower, with achange in altitude without throttle adjustmentby the pilot. Fuel flow in the MANUAL mode,however, is not temperature compensated.Because of this, there may be temperatures atwhich maximum fuel flow in MANUAL modewill not be sufficient to achieve TakeoffPower.

Table 1-1: Time to Power ChangeENGINE POWER AT TIME OF FADEC FAILURE

DESIRED POWER AS SELECTED BY THROTTLE POSITION

APPROX. TIME TO DETECT POWER CHANGE DURING TRANSITION TO MANUAL

LOW POWER HIGHER POWER 2.0 SECONDSLOW POWER LOWER POWER 1.0 SECONDHIGH POWER HIGHER POWER 7.0 SECONDSHIGH POWER LOWER POWER 0.1 SECOND


31 JAN 2007—Rev. 2———1-43

NOTE

In the event engine (NP) overspeedsystem is activated during transitionto or operation in MANUAL mode, thecontrol system is designed to keepthe engine running. Engine mayoscillate between 112.5 and 118.5%NP until corrective action is takenwi th thro t t le and co l lec t ive(paragraph 1-12-N).

3. FADEC MODE switch – Depress onetime.

Depressing the FADEC mode switch onetime will mute the FADEC FAIL warninghorn (chime tone). As the transition toMANUAL mode was init iated by theFADEC system, this step should not beaccompl ished unt i l p i lo t is f i rmlyestablished in MANUAL control.

This will allow pilot to keep hands onflight controls during the transition toMANUAL mode.

If the FADEC ECU is operational, it willt rack HMU opera t ion , per formdiagnostics, monitor engine functions,and provide overspeed limiting for bothNP and NG. Surge detec t ion andavoidance will not be available. If anengine surge is encountered, decreasethe throttle until the surge conditionclears, then slowly increase the throttleto the desired power level. Rapid powerchanges should be avoided.


Applicable maintenance action will berequired prior to next flight.

5. Normal shutdown if possible.

If normal shutdown can not be completedby rolling throttle to closed position, fuelshutoff valve can be positioned to off.

1-12-J. CATEGORY 2 — FADECDEGRADED

FADEC DEGRADED faults represent a loss ofsome feature of the FADEC system which maycause a degradation in performance. Thismay result in NR droop, NR lag, or reducedmaximum power capability. These faults willbe displayed immediately when detected bythe ECU. Operations should be continued inAUTO mode and helicopter is to be flownsmooth ly and nonaggress ive ly. Inconjunction with the FADEC DEGRADEDlight, the RESTART FAULT light may alsoactivate under certain fault conditions.

With FADEC Software Version 5.356 installed,the reversionary (backup) governor will beactivated under certain fault conditions,which will also allow operations in a degradedmode (paragraph 1-12-M).

Applicable maintenance action wil l berequired prior to next flight.

1-12-K. CATEGORY 3 — FADEC FAULT

FADEC FAULT indicates that PMA and/or MGT,NP, or NG automatic limiting circuit(s) may notbe functional. In conjunction with activationof the FADEC FAULT light, the RESTARTFAULT light may also activate under certainfau l t condi t ions . These fau l ts wi l l bedisplayed immediately when detected by theECU. Operations should continue in AUTOmode. If both lights (FADEC FAULT andRESTART FAULT) are i l luminated, thisindicates the MGT automatic limiting circuit(905°C in flight), may not be functional. Thepi lo t should fo l low the appropr ia teprocedures as set out in the BHT-407-FM-1.

Applicable maintenance action wil l berequired prior to next flight.

1-12-L. CATEGORY 4 — RESTARTFAULT

RESTART FAULT indicates a subsequentautomatic engine start may not be possible.


1-44———Rev. 2—31 JAN 2007

The fault does not require immediate actionby the p i lo t and shou ld not a f fec tper formance of the he l icop ter. I t i srecommended that the pilot plan the landingsi te accord ingly. These fau l ts wi l l bedisplayed immediately when detected by theECU and displayed as RESTART FAULT. Donot a t tempt a subsequent s ta r t un t i lapplicable maintenance action has beencompleted.

If the engine shutdown procedures are notproperly followed, the MANUAL mode pistonsmay begin to engage during the shutdown.The FADEC may then be unable to prevent ahot start on the next start and will indicate aRESTART FAULT to warn the pilot. An HMUmanual piston parking procedure will berequired, as described in paragraph 1-12-Pand the Rolls-Royce 250-C47B Operationsand Maintenance Manual.

1-12-M. FADEC REVERSIONARYGOVERNOR (FADEC SOFTWAREVERSION 5.356)

The FADEC Reversionary (backup) Governorconsists primarily of a backup channel whichis contained in the ECU and is isolated fromthe primary governor by a firewall for EMI. Itprov ides bas ic power turb ine speedgoverning in the event of a hard faultoccurring in the primary governor. Failure ofthe FADEC into Reversionary Governor modeis indicated by the i l luminat ion of thefollowing three lights: FADEC FAULT, FADECDEGRADE, and RESTART FAULT. This type offailure causes a degradation in performanceand can cause NR droop, NR lag, or reducedmaximum power capabil ity. Operationsshould be continued in AUTO mode andhelicopter is to be flown smoothly andnon-aggressively. Applicable maintenanceaction will be required prior to next flight(paragraph 1-12-R).

1-12-N. ENGINE OVERSPEEDPROTECTION

NP overspeed limiting is available in both theAUTO and MANUAL modes by independentanalog circuits integral to the ECU. NGoverspeed limiting is available in the AUTOand MANUAL modes through software controlin the ECU. In the event of a FADEC FAILURE,it is possible that NG overspeed protectionwill not be available.

The FADEC ECU continuously monitors forNG and NP, or NP versus torque (Q) (5.202FADEC software) overspeed conditions inboth AUTO and MANUAL Mode.

Activation of the ENGINE OVSPD warningl ight w i l l occur in the event o f an NPoverspeed, NG overspeed, or if NP versustorque (5.202 FADEC software) is above thecontinuous limit (102.1% NP at 100% torque to108.6% NP at 0% torque). With 5.356 FADECsoftware, the ENGINE OVSPD warning lightwill illuminate when NP reaches or exceeds102.1%, regardless of the torque value. Thelight will also momentarily illuminate duringthe overspeed system test when theoverspeed solenoid valve closes.

If the ENGINE OVSPD light is activated duringengine operation due to an exceedance, andthe value has been recorded by the ECU, thepilot will be provided with a maintenanceadvisory on shutdown in the form of a FADECDEGRADED light. The FADEC DEGRADEDlight will illuminate when NG speed decaysbelow 9.5%. If the pilot fails to recognizeillumination of the FADEC DEGRADED lighton shutdown, it will be illuminated the nexttime electrical power is applied following theFADEC system self test. When the FADECDEGRADED l ight is i l lumina ted as amaintenance adv isory, maintenanceinvestigation is required prior to further flight.Peak values of exceedances are located onthe Engine History Data page of the EMC-35AMaintenance Terminal.


31 JAN 2007—Rev. 2———1-45

If limits are exceeded, refer to 250-C47BSeries Overspeed Limits in the Rolls-RoyceOperation and Maintenance Manual and theBHT-407-MM-2, Chapter 5.

• NP OVERSPEED

When the engine reaches 118.5 ±1% NP,overspeed limiting will occur. The analogoverspeed limiting feature will activatethe overspeed solenoid valve, whichreduces fuel to the engine to a minimumflow condition (sub-idle value of 34 to 45pph). The minimum fuel flow increasesthe likelihood of the engine remainingrunning and recover ing f rom theoverspeed. Once the NP speed drops to112.5 ±2%, the overspeed solenoid valvewill be deactivated and fuel flow willreturn to its previously commandedvalue.

In the event the overspeed cannot becontrolled after fuel flow is reintroduced,the overspeed limiting feature will controlthe overspeed between the activation trippo in t o f 118 .5 ±1% N P and thedeactivation point of 112.5 ±2% NP. If thisoccurs, attempt to control engine androtor speed with throttle and collective.Refer to the ENGINE OVERSPEEDprocedure in the Flight Manual.

• NG OVERSPEED

In AUTO mode, a software implementedoverspeed system is provided. Shouldthe software detect an NG overspeed, theprotection feature will be activated. Inaddition, if the ECU has not failed, NGoverspeed protection will be available inMANUAL mode.

When the engine reaches 110 ±1% NG,the ENGINE OVSPD warning light willilluminate and overspeed limiting willoccur. The software controlled overspeedl imi t ing fea tu re w i l l ac t iva te theoverspeed solenoid valve which reducesfuel to the engine to a minimum flow

condition (sub-idle value of 34 to 45 pph).The minimum fuel flow increases thel ikel ihood of the engine remainingrunning and recover ing f rom theoverspeed. Once the NG speed drops to107 ±1%, the overspeed solenoid valvewill be deactivated and fuel flow willreturn to its previously commandedvalue.

In the event the overspeed cannot becontrolled after fuel flow is reintroduced,the overspeed limiting feature will controlthe overspeed between the activationpoint of 110 ±1% NG and the deactivationpoint of 107 ±1% NG. If this occurs,attempt to control engine and rotorspeed with throttle and collective. Referto the ENGINE OVERSPEED procedure inthe BHT-407-FM-1.

• OVERSPEED SYSTEM SHUTDOWN TEST

Functionality of the overspeed system ischecked during FADEC power up andthereafter continuously by the ECU.Operation of the overspeed solenoid ischecked periodically by the pilot throughthe use of the OVERSPEED SHUTDOWNtest procedure.

The OVERSPEED SHUTDOWN testprocedure will shut down the engine onlyif collective pitch is below 10%, throttleposition is at idle, NG is between 60 to66% and NP is less than 75%. TheOVERSPEED test but ton must bepressed and held for a minimum of 1.0second but not more than 10.0 seconds.Once the test button is released, theOVERSPEED test is completed asfollows. The FADEC ECU signals theoverspeed solenoid valve to close andthe ENGINE OVSPD light to come on.Once the FADEC ECU senses an NGdecrease greate r than 0 .5%, theoverspeed solenoid valve is opened, theENGINE OVSPD light goes off, and theengine is shut down by FADEC ECUactivation of the hot start abort feature.


1-46———Rev. 2—31 JAN 2007

If the overspeed test is unsuccessful, theengine will continue to operate at idlepower, the FADEC FAULT caution lightwill illuminate, and a normal shutdownprocedure must be carried out.

1-12-O. ENGINE SHUTDOWN

Pilot control of engine speed from 100% NR/NP to idle in AUTO mode is controlled throughthro t t le movement . As the thro t t le ispositioned from the detented FLY position toidle, electrical signals are sent to the ECUfrom the HMU – PLA potentiometer. Thesesignals dictate the amount of authority theECU has to control maximum fuel flow (NGlimiting), and in turn, engine speed. Therefore,as throttle is decreased, the maximum fuelflow that can be delivered to the engine isreduced by positioning the fuel meteringvalve to control engine NG speed/power.

In the unlikely event that a system faultoccurs which does not allow a reduction inengine speed by positioning the throttle toidle, complete the 2 minute cool-down at100% flat pitch. After the 2 minute cool-down,the engine is to be shutdown by rolling thethrottle to the CLOSED position.

Pilot control of engine speed from 100% NR/NP to NG idle in MANUAL mode is controlledhydromechanica l ly through th rot t lemovement. Idle speed in MANUAL mode maynot stabilize at 63 ±1% NG. If this occurs,maintain idle speed at 63 ±1% NG with throttle.

Following the appropriate cool down period atidle, the engine may be shut down in eitherthe AUTO or MANUAL mode by positioningthe throttle to cutoff. This will close themechanical fuel shutoff valve within the HMU.

Do not reposition the throttle out of cutoffunless NG has decayed to zero. If the throttleis positioned out of cutoff prior to the NGspeed decreasing through 9.5%, the FADECIN-FLIGHT restart logic will introduce fuel andactivate the igniter. This can cause a relightand possible over temperature condition(paragraph 1-18).

If relight occurs, the pilot must immediatelyposition the throttle to the closed positionand activate the starter.

Additionally, the pilot must also allow NG todecay to 0% in AUTO mode pr ior topositioning the battery switch to OFF. If thisprocedure is not followed, the MANUAL modepistons may move hydromechanically afterelectrical power is removed. This may cause aRESTART FAULT the next time power isapplied to the FADEC ECU and a HMU manualpiston parking procedure will be required perparagraph 1-12-P or as described in theRo l ls -Royce 250-C47B Operat ion andMaintenance Manual.

1-12-P. HMU MANUAL PISTON PARKINGPROCEDURE

Starting with the HMU MANUAL mode pistonsin the wrong position may result in a hot startof the engine. When the pilot is not certain ofthe position of the pistons, or has received aMaintenance Advisory that the pistons are outof position, the following procedure willassure the pistons are in the correct position(fully retracted) for engine starting.

6. Position throttle to cutoff.

7. Pull IGNITER circuit breaker.

8. BATT — ON.

9. Power up check — Complete.

10. FADEC Mode switch — MANUAL.

11. Motor the engine (with throttle incutoff) for 10 seconds.

12. Wait for NG to decay to 0%.

13. FADEC Mode switch — AUTO.

14. Motor the engine (with throttle incutoff) for an additional 10 seconds.

15. Wait for NG to decay to 0%.

16. BATT — OFF.

17. Push in IGNITER circuit breaker.


31 JAN 2007—Rev. 2———1-47

Continue prestart checklist.

1-12-Q. CATEGORY 5 — MAINTENANCEADVISORY, FADEC SYSTEMFAULTS — ENGINE SHUTDOWN

Maintenance Advisory Faults are thosedetected by the ECU that are consideredminor in nature and are not communicated tothe cockpit with the engine operating. TheFADEC DEGRADED light serves as themaintenance advisory light. This light will beilluminated, upon engine shutdown, if anyfault or exceedance has been detected duringthe last engine run or if a current fault exists.This will indicate that maintenance action isrequired prior to the next flight.

Maintenance advisory faults will displayduring shutdown when the throttle is placedin the CLOSED position and the NG speeddecays below 9.5%. If the pilot misses themaintenance advisory on shutdown, it willilluminate at the next application of electricalpower.

1-12-R. CHECKING FADEC FAULTCODES

As stated previously, faults can be displayedimmediately via a FADEC FAIL, FADECMANUAL, FADEC DEGRADED, FADEC FAULT,RESTART FAULT or by a combinat ionof these lights. Maintenance Advisory faultswill be displayed on shutdown via the FADECDEGRADED light. In addition, faults are alsoused to identify exceedances.

Regardless if the fault light(s) were displayedin-flight or at shutdown, maintenance actionis required prior to further flight. Refer to theBHT-407-MM-9, Chapter 76 and Rolls-Royce250-C47B Operation and Maintenance Manual.

The preferred method of determining FADECfaults or exceedances is with the GoodrichEMC-35A Main tenance Termina l . TheEMC-35A Maintenance Terminal (Windowsversion) is capable of providing informationon Current Faults, Last Engine Run Faults,

Accumulated Faults (Fault History screen),and NP overspeed exceedance information(Engine H is tory screen) . Refer to theGoodrich Maintenance Terminal User Guidefor operating instructions.

If a maintenance terminal is not available,identify faults or exceedances using theMaintenance Mode feature of the FADECsystem. This feature allows operators todetermine faults through a sequence offlashing light displays on the cockpit cautionpanel (paragraph 1-12-R-1).

1-12-R-1. ENGINE RUN FAULT CODES —PROCEDURE FOR VIEWING

The caution panel fault display may beoperated as follows:

NOTE

Displayed faults may be LASTENGINE RUN or CURRENT faults.Following this procedure, refer toparagraph 1-12-R-2 to determine iffaults are current.

1. Engine must be shut down and theFADEC MODE switch positioned toMANUAL. Place the collective fulldown (below 10%) and the throttle inthe cutoff position.

NOTE

If the throttle or collective is movedduring the above procedure or theFADEC MODE switch is positioned toAUTO, the FADEC ECU will exit thefault code reporting mode.

2. Depress and release the FADEC ECUmaintenance button on the left handside of the lower pedestal to enter thefault code reporting mode.

3. FADEC DEGRADED, FADEC FAULT,and the RESTART FAULT lights willsimultaneously flash five times to


1-48———Rev. 2—31 JAN 2007

indicate that maintenance mode hasbeen entered by the ECU.

4. Depress and release the FADEC ECUmaintenance button. If a fault ispresent, it will be displayed by aspecified number of FADECDEGRADED caution panel lightsegment flashes.

5. Depress and release the FADEC/ECUmaintenance button to flash the nextfault code.

6. Steady illumination of the FADECDEGRADED caution panel lightsegment indicates that no otherfaults exist for this light.

7. Continue to depress and release theFADEC/ECU maintenance button tostep through the FADEC FAULT andRESTART FAULT caution panel lightsegments, as above. This willdetermine if fault codes exist forthese segments.

8. If no fault code exists for the selectedcaution panel segment, the cautionlight will illuminate continuouslywhen the FADEC/ECU maintenancebutton is released.

9. When interrogation is complete, thenext push of the FADECmaintenance button will cause theFADEC DEGRADED, FADEC FAULT,and RESTART FAULT caution lightsegments to flash simultaneouslyfive times and then extinguish. Thisindicates that the FADEC ECU hasexited the maintenance mode.

NOTE

Refe r to Tab le 1 -2 for FADECSoftware Version 5.202 and Table 1-3for FADEC Software Version 5.356

(Reversionary Governor) Fault CodeDisplays.

10. To determine fault description andmaintenance message code(s) fromcaution panel flashing display, referto Table 1-2 or Table 1-3.

1-12-R-2. FADEC FAULT CODES —PROCEDURE TO DETERMINELAST ENGINE RUN FAULTSFROM CURRENT FAULTS

The procedure to determine if the fault codesdisplayed are LAST ENGINE RUN faults orCURRENT faults may be accomplished byperforming the steps listed in paragraph1-12-R-1 with the throttle in the IDLE position.

With the throttle positioned to IDLE, anyFADEC fault code which is displayed will be acurrent fault.

In addition, if no FADEC related lights aredisplayed on the caution, warning, advisorypanel with electrical power applied, theFADEC in AUTO mode, and the throttlepositioned to idle, no current faults exist.

1-12-R-3. FAULT CODE CHARTS — USE OF

If fault codes have been displayed via thecaution panel, refer to Table 1-2 or Table 1-3for specific information. This information canbe used in conjunct ion w i th theBHT-407-MM-9 , Chapte r 76 and theRolls-Royce Operation and MaintenanceManual 250-C47B. These publications containthe data to determine the maintenance actionrequired prior to further flight.

For specific fault information displayed on theEMC-35A Maintenance Terminal, refer to theRo l ls -Royce 250-C47B Operat ion andMaintenance Manual to determine themaintenance action required prior to furtherflight.


31 JAN 2007—Rev. 2———1-49

Table 1-2: 250-C47B FADEC Software Version 5.202 Fault Code Display

STATUS MESSAGE FAULT DESCRIPTION MAINTENANCE MESSAGE CODE

FADEC DEGRADED cockpit lamp flashes 1 time.

ECU Failure has occurred. AD12bitFlt, AD8bitFlt, ECUOTFlt, GainFlt, HLRfFLt, OffsFlt, PROMFlt, PW10Flt,RAMFlt, V15Flt, V5Flt, WDTFlt, CJCFlt, OrDiodeFlt, BacCompFlt, EEPROMFlt, ForCompFlt, P1Flt, SWIntFlt, TestCelFlt, UARTFlt, UUIntFlt, WDTOutFlt, OSVFlt, NpOSFlt, OR28Flt, WDTTimeOut

FADEC DEGRADED cockpit lamp flashes 2 times.

This Status Message is not used. Verify that message is indicated by lamp.

Not Used


NP-Q Exceedance NpQExLmAdv



Not Used


MGT indication failure MGTFlt



Not Used


Failure to control HMU – Auto/Manual Solenoid

AMSolFlt


CIT temperature indication failure

T1AFlt, T1BFlt, or T1ABFlt


Metering Valve not in start position

OpenMvFlg


Starter Relay Interface StrFlt


NR Sensor – Rotor decay anticipation

NrFlt


Incorrect Overspeed Test Switch indication

OSTstSwFlt


1-50———Rev. 2—31 JAN 2007

FADEC FAULT cockpit lamp flashes 1 time.

HMU – Failure to control fuel flow

WfLimFlag

FADEC FAULT cockpit lamp flashes 2 times.


Not Used


NP-Q Run Limit advisory NpQRnLmAdv


HMU Metering Valve Potentiometer

WfMvFlt or WfStFlt


Collective Pitch Potentiometer indication failure

CPFlt


Failure to control HMU (Stepper motor)

StepCntFlt, SMFlt, or AMSolFlt



Not Used


Failure to control HMU (Overspeed Solenoid)

OSFlt


Engine Surge event SgFlag


NG speed indication failure Ng1Flt, Ng2Flt, or Ng12Flt


Airframe Power Supply failure

AF28Flt


Engine Overspeed OSFlag

RESTART FAULT cockpit lamp flashes 1 time.


Not Used

RESTART FAULT cockpit lamp flashes 2 times.

TMOP Sensor – Torque indication failure

QFlt

Table 1-2: 250-C47B FADEC Software Version 5.202 Fault Code Display (Cont)



31 JAN 2007—Rev. 2———1-51



Not Used



Not Used


Failure in HMU (PLA Potentiometer) position reading

PLA1Flt, PLA2Flt, PLA12Flt or PLARfFlt


PMA power supply failure Al28Flt


Failure to control HMU (Hot Start Abort Solenoid)

StSFlt or StSIFlt



Not Used


Failure to control Ignition Relay Interface

IgnFlt or IgnIFlt


NP speed indication failure Np1Flt, Np2Flt, or Np12Flt


Incorrect Auto/Manual switch indication

AMSwFlt


Quiet Mode switch fault QMSwFlt

FADEC DEGRADE, FADEC FAULT and RESTART FAULT cockpit lamps on steady.

All faults have been displayed.

Table 1-2: 250-C47B FADEC Software Version 5.202 Fault Code Display (Cont)



1-52———Rev. 2—31 JAN 2007

Table 1-3: 250-C47B FADEC Software Version 5.356 (Reversionary Governor)Fault Code Display


FADEC DEGRADED cockpit lamp flashes 1 time.

Primary Governor Failed AD12bitFlt, AD8bitFlt, ECUOTFlt, GainFlt, HLRfFLt, OffsFlt, PROMFlt, PW10Flt, RAMFlt, V15Flt, V5Flt, WDTFlt, CJCFlt, OrDiodeFlt, BacCompFlt, EEPROMFlt, ForCompFlt, P1Flt, SWIntFlt, TestCelFlt, UARTFlt, UUIntFlt, OSVFlt, NpOSFlt, OR28Flt, WDTTimeOut, ARINCFlt, ARINCHWFlt, SWCfgFlt, RGSDFlt,QRawFlt, T1BRawFlt, ESWRGFlt, ESW2RGFlt, ESW3RGFlt, ESW4RGFlt, ESW5RGFlt, SWConfigRGFlt, PwrRstFlt, RGSelSwFlt or NDOTWRCdRGFlt


Reversionary Governor Failed

AD10bitFltRG, PROMFltRG, RAMFltRG, RGOTFltRG, SWConfigFltRG, V10FltRG, V15nFltRG, V15pFltRG, V5qFltRG, WDTTimeOutRG, ARINCHdFltRG, ARINCFltRG, ARINCHWFltRG, BacCompFltRG, ForCompFltRG, Or28FltRG, OrDiodeFltRG, PW10LoFltRG, RGTempFltRG, SPITempFltRG, UARTFltRG, WDTFltRG, PGSDHdFltRG, PGSDFltRG, WfCorrPGHdFltRG, WfCorrPGFltRG, EngRnCtPGFltRG, EngRnTmPGFltRG, ESWPGFltRG, NpIncPGFltRG, Np2RawPGFltRG, P1RawPGFltRG, T1ARawPGFltRG, SwPwrFltRG


NP Exceedance NpLmTOut


31 JAN 2007—Rev. 2———1-53


Reversionary Governor did not govern when Primary Governor failed.

ECUGovFltRG


MGT indication failure MGTFlt



Not Used


Failure to control Auto/Manual Solenoid

AMSolFlt, AMSolFltRG


CIT temperature indication failure

T1AFlt, T1ABFlt, T1BFltRG or T1DFltRG


Metering Valve is not in start position.

OpenMvFlg


Starter Relay Interface StrFlt


NR Sensor – Rotor decay anticipation

NrFlt


Incorrect Overspeed Test switch indication

OSTstSwFlt

FADEC FAULT cockpit lamp flashes 1 time.

HMU – Failure to control fuel flow

WfLimFlag, WfLimFlagRG



Not Used


NP Run Limit NpRLmTOut

Table 1-3: 250-C47B FADEC Software Version 5.356 (Reversionary Governor)Fault Code Display (Cont)



1-54———Rev. 2—31 JAN 2007


Failure in HMU Metering Valve Potentiometer reading

WfMvFlt, WfStFlt or WfStFltRG


Collective Pitch Potentiometer indication failure

CPFlt


Failure to control HMU (Stepper motor)

StepCntFlt, SmFlt, AMSolFlt or SmFltRG



Not Used


Failure to control HMU (Overspeed Solenoid)

OSFlt


Engine Surge event SgFlag


NG speed indication failure Ng1Flt, Ng2Flt, Ng12Flt or Ng1FltRG


Airframe Power failure AF28Flt or AF28FltRG


Engine Overspeed OSFlag or OSEventLmpRG

RESTART FAULT cockpit lamp flashes 1 time.


Not Used


TMOP Sensor – Torque indication failure

QFltRG



Not Used



Not Used




31 JAN 2007—Rev. 2———1-55

1-12-S. CLEARING FADEC FAULTCODES

Faults/exceedances are not to be erasedunless appropriate maintenance actions havebeen carried out in accordance with theBHT-407-MM and Rolls-Royce 250-C47BOperations and Maintenance Manual. Do notattempt to clear any fault or exceedance whilethe engine is operating.

If maintenance actions have been conducteddue to a recorded exceedance or to correct acurrent or last engine run fault, it must be

ensured that no FADEC system lights areilluminated when the throttle is positioned toIDLE for the next start attempt. If a FADECrelated light is illuminated with the throttlepositioned to IDLE, a current fault exists andfurther maintenance action is required.

1-12-S-1. CURRENT FAULTS

Current faults may be cleared by performing apower reset (battery switch OFF/ON). If fault isno longer detected, associated FADECDEGRADED light will be extinguished.


Failure in HMU (PLA Potentiometer) reading

PLA1Flt, PLA2Flt, PLA12Flt or PLARfFlt


PMA power supply Al28Flt or Al28FltRG


Failure to control HMU (Hot Start Abort Solenoid)

StSFlt, StSIFlt, StSFltRG or StSIFltRG



Not Used


Failure to control Ignition Relay Interface

IgnFlt or IgnIFlt


NP speed indication failure Np1Flt, Np2Flt, NpDFlt, Np12Flt, NpDFltRG or Np1FltRG


Incorrect Auto/Manual switch indication

AMSwFlt or AMSwFltRG


Quiet Mode switch fault QMSwFlt

FADEC DEGRADE, FADEC FAULT, and RESTART FAULT cockpit lamps on steady.

All faults have been displayed.




1-56———Rev. 2—31 JAN 2007

1-12-S-2. LAST ENGINE RUN FAULTS/EXCEEDANCES

LAST ENGINE RUN faults or exceedancesmay be cleared by performing a successfulengine start or with use of the EMC-35AMaintenance Terminal.

To e rase LAST ENGINE RUN fau l ts /exceedances , re fer to the EMC-35AMaintenance Terminal Users Guide foroperating instructions.

1-12-S-3. ACCUMULATED FAULTS/EXCEEDANCES

Accumulated faults or exceedances may onlybe c leared w i th use o f the EMC-35AMaintenance Terminal.

NP exceedance values may only be clearedfrom engine history data page with use of theEMC-35A Maintenance Terminal.

1-12-T. FADEC TRAINING IN MANUALMODE

Prior to actual flight training in the MANUALmode, an operational understanding of theFADEC system along with a sound knowledgeof emergency procedures is required. It isrecommended that t ra in ing be f i rs taccomplished in the helicopter in a cockpitprocedural environment (engine not running)(paragraph 1-12-U). This can provide a visualsimulation of the horn and lights associatedwith a FADEC failure direct to MANUAL modeand familiarize the pilot with the requiredcockpit actions. This should be followed by atakeoff/hover/circuit and landing in MANUALmode which will allow the pilot to becomefamiliar with the required manipulation of thethrottle and controls (paragraph 1-12-V).

Once the pilot is comfortable with flight inMANUAL mode, simulated FADEC failureemergency procedures can be carried out inflight (paragraph 1-12-W).

1-12-U. COCKPIT PROCEDURALTRAINING ON GROUND (ENGINENOT RUNNING)

1. Pull START and IGNTR circuitbreakers and ensure fuel valveswitch is positioned to OFF.

2. Connect external power source.

3. Allow instruments and FADEC tocomplete self test with FADEC MODEswitch in AUTO.

4. Position throttle to FLY detent andcollective to approximate cruiseflight setting.

5. Simulate FADEC FAILURE by pullingFADEC circuit breaker. This willsimulate a failure direct to MANUALmode. FADEC FAIL warning horn willactivate along with illumination ofthe FADEC FAIL and FADECMANUAL caution panel lights.FADEC MODE switch will illuminateMAN.

6. Carry out the appropriateBHT-407-FM-1 emergency responseprocedure. (Depressing the FADECMODE switch will silence the horn.)

7. Push in FADEC circuit breaker andset FADEC MODE switch to AUTO.

8. Repeat procedure until it isunderstood.

9. Disconnect external power source,position throttle to cutoff and push inSTART and IGNTR circuit breakers.

1-12-V. FLIGHT TRAINING IN MANUALMODE

In MANUAL mode, the following is applicable:

• Switching to MANUAL mode at IDLE(63 ±1% NG) may result in change inNG.

• Igniter operates continuously.


31 JAN 2007—Rev. 2———1-57

• AUTO RELIGHT, FADEC MANUALcaution panel lights illuminate.

• FADEC MODE switch indicates MAN.

• FADEC ECU remains operational.However, surge protection andavoidance logic is not available. In theevent of a FADEC FAILURE whiletraining in MANUAL mode, the FADECFAIL warning light will illuminate.FADEC FAIL horn will not sound.Remain in MANUAL mode and land assoon as practical.

• Maximum Continuous Power isavailable for all ambient conditions.Takeoff Power may not be available.

1. Perform START procedure andSYSTEMS CHECKS in AUTO mode.

2. At idle (63 ±1% NG), depress FADECMODE switch to transition toMANUAL mode.

NOTE

Transition back to AUTO mode canbe made at any time by depressingFADEC MODE switch to AUTO. Uponselecting AUTO mode, an enginepower transient may be experiencedas the FADEC ECU matches enginepower to rotor load. Ensure throttle ispositioned to FLY detent positionfollowing selection of AUTO mode.

3. Manipulate throttle on ground tobecome familiar with MANUALcontrol.

4. Increase throttle to maintain 95 to100% NR/NP. Manipulate throttle/collective and lift into hover.

5. When comfortable with manipulationof throttle and flight controls in

hover, transition into forward flightand conduct circuits to touchdown.

1-12-W. SIMULATED FADEC FAILURETRAINING (IN-FLIGHT)

When comfortable with flying circuits to touchdown in MANUAL mode, transitions betweenAUTO and MANUAL may be conducted inflight.

Simulation can be accomplished by activatingthe FADEC FAIL horn by pushing the FADECFAIL horn test button. Although the applicableFADEC caution panel l ights cannot beactivated, the pilot should respond to theFADEC FAIL horn and carry ou t theBHT-407-FM-1 emergency responseprocedure. Once the throttle bezel position ismatched to the actual NG indication, pilotshall position the FADEC MODE switch toMANUAL.

Following simulated FADEC FAILURE trainingin flight, ensure FADEC MODE switch ispositioned to AUTO prior to shutdown. Thiswill ensure the manual pistons in the HMU areparked, and a subsequent start in AUTO canbe carried out without maintenance action.

1-13. COMBINED ENGINE FILTERASSEMBLY (CEFA)

The CEFA (Figure 1-13) provides both fuel andscavenge lube filtration within a single filterassembly which consists of a fuel filter bowl,fuel bypass valve, fuel differential pressureindicator, manifold assembly, a disposablefuel filter element, a lube filter bowl, lubebypass valve, lube differential pressureindicator, and a disposable lube filter element.The CEFA is located on the lower left handportion of the power and accessory gearboxassembly.


1-58———Rev. 2—31 JAN 2007

Figure 1-13. Power Plant Components (Sheet 1 of 2)407_MD_01_0011

7

Ignition exciter box (Ref)

Torque meter oil pressure port (Ref)

Engine oil outlet port (Ref)

Engine oil inlet port (Ref)

Engine oil pressure port (Ref)

1.

2.

3.

4.

5.

Engine intake (Ref)

Shroud bleed air manifold connection (Ref)

Oil filter assembly

Bleed air valve (acceleration)

CEFA fuel filter impending bypass button

6.

7.

8.

9.

10.

6

5

4

3

2

1

10

9

COMPRESSOR TURBINE COMBUSTION SECTION

POWER ACCESSORIES GEAR BOX

TYPICAL

8


31 JAN 2007—Rev. 2———1-59

Figure 1-13. Power Plant Components (Sheet 2 of 2)407_MD_01_0012

22

Fuel nozzle (Ref)

Ignitor

MGT terminal block

NP monopole pickup

NG monopole pickup

Anti-ice solenoid valve

Chip plug forward

11.

12.

13.

14.

15.

16.

17.

Starter-generator

Hydro mechanical unit (HMU)

Scavenge oil filter and impending bypass button

Combined engine filter assembly (CEFA)

Fuel filter

Permanent magnet alternator (PMA)

Hose assembly, combustion chamber drain (Ref)

18.

19.

20.

21.

22.

23.

24.

20

24

1819

11

23

17

12

16

14

15

13

TYPICAL

10

21


1-60———Rev. 2—31 JAN 2007

The bypass valves for both the fuel andscavenge lube filters allow flow to go into abypass condition if excessive differentialpressure occurs across the filter elements.The differential pressure indicators provide avisual signal when differential pressureacross e i ther e lement exceeds apredetermined value indicating that theelement is dirty and requires replacement.Both indicators are reset manually followingreplacement.

The fuel bypass valve works in conjunctionwith an impending bypass indicator. Whendifferential pressure across the fuel filter risesto a pressure slightly less than the pressureat which the bypass valve actually cracks, ared button extends, signalling that a bypass isabout to occur. The indicator actuates at 2.1to 2.9 PSI and the bypass valve cracks at 3.4PSI minimum. The bypass valve seats whenthe pressure drops to 3.0 PSI. The fuel filterimpending bypass indicator is located on thelower outboard side of the manifold assembly.

The scavenge lube filter bypass indicatorvisual ly s ignals an impending bypasscondition by extending a red button when thedifferential pressure across the filter isbetween 8.8 to 10.8 PSI. The lube bypassindicator assembly is equipped with a thermallockout for temperatures below 43°C ±17°C(110°F ±15°F). The lube indicator assembly ismounted at the bottom (aft end) of the lubebowl.

If the fuel impending bypass indicator or lubefilter bypass indicator extends, ensureappropriate Rolls-Royce engine maintenanceprocedures are carried out prior to furtherflight.

1-14. POWER PLANT IGNITIONSYSTEM

The ignition system consists of a harness, asolid-state low tension capacitor dischargeignition exciter, a spark igniter lead, and ashunted sur face-gap spark ign i ter(Figure 1-13).

The ignition system transfers energy to thecombustible fuel mixture. Energy is providedin the fo rm of h igh tempera tu re -h ighamperage arcs at the spark igniter gaps.These arcs ignite the fuel/air mixture. Theignition exciter is only required during thestarting cycle since the combustion processis continuous. Once ignition takes place, theflame in the combustion liner acts as theignition agent for the fuel/air mixture.

Dur ing FADEC manua l opera t ion , theautomatic auto relight circuit is deactivatedand the igniter is turned on at all times at gasproducer (NG) speeds above 55% to reducethe possibility of engine flameout.

1-15. POWER PLANTTEMPERATUREMEASUREMENT SYSTEM

The temperature measurement systemconsists of four chromel-alumel singlejunction thermocouples in the gas producerturbine outlet and an associated integralharness . The vo l tages o f the fourthermocouples are electrically averaged inthe assembly. The harness terminates at aterminal block on the aft side of the horizontalfire shield. The engine electrical harness, thehelicopter MGT indicator wiring, and theFADEC MGT input wiring, all connect to thisterminal block (Figure 1-13).

1-16. POWER PLANTCOMPRESSOR BLEED AIRSYSTEM

The compressor bleed air system permitsrapid engine response. The system consistsof a bleed control valve located on the frontface of the scroll and an inducer bleedmani fo ld which encases the s lo t tedcompressor shroud housing.

The bleed control valve is open duringstarting and ground idle operation, andremains open until a predetermined pressureratio is obtained. At the predetermined


31 JAN 2007—Rev. 2———1-61

pressure ratio, the valve begins to modulatefrom the open to the closed position.

The inducer b leed d ischarges a i r toatmosphere at engine idle speed. At higherpower settings, flow changes from bleed tointake air.

1-17. ENGINE OIL SYSTEM

The engine incorporates a dry sump oilsystem (Figure 1-14) with an externallymounted supply tank and oil cooler locatedon the top aft section of the fuselage andenclosed by the aft fairing. Oil is suppliedfrom the tank to gear type pressure andscavenge pumps mounted within the engineaccessory drive gearbox. A spur gear type oilpump assembly, consisting of one pressureelement and four scavenge elements ismounted within.

The oil filter assembly, consisting of an oilfi lter, fi l ter bypass valve, and pressureregulating valve, is located in the top left-handside of the gearbox (Figure 1-13). A checkvalve, located between the filter package andthe accessory gearbox, prevents oil fromdraining into the engine from the helicoptertank when the engine is not in operation.

Additional scavenge lube filtration is providedwithin the combined engine filter assembly(CEFA) (paragraph 1-13).

Indicating type magnetic chip detectors areinstalled at the bottom of the gearbox and atthe engine oil outlet connection. All engine oilsystem lines and connections are internalexcept the pressure and scavenge lines to thefront compressor bearing and to the bearingsin the gas producer and power turbinesupports.

The system is designed to furnish adequatelubrication, scavenging, and cooling asneeded to the bearings, splines, and gearsregardless of the helicopter attitude oraltitude. Jet lubrication is provided to all

compressor, gas producer turbine, and powerturbine rotor bearings, and to the bearingsand gear meshes of the power turbine geartrain with the exception of the power outputshaft bearings. The power output shaftbearings and all other gears and bearings arelubricated by oil mist.

NOTE

If helicopter engine has been shutdown for more than 15 minutes,scavenge oil could have drained intogearbox. Dry motor run engine for 30seconds before checking oil level. Ifnot accomplished, a false highengine o i l consumpt ion ra teindication or overfilling of oil tankcould result.

The approximate capacity of the engine oiltank is 1.5 US gallons, and the oil level ischecked by means of a sight gauge mountedon the left side of the tank. Viewing access tothe sight gauge is provided by a cutout in thecowling. The oil cooler is mounted on top ofthe duct on the oil cooler blower.

1-18. ENGINE AUTO RELIGHT

• AUTO MODE — AUTO RELIGHT

In AUTO mode, the FADEC is capable ofdetecting an engine flameout by measuringan NG deceleration rate greater than thepredetermined flameout boundary rate. If aflameout is detected, the ENGINE OUTwarning light and horn will be activated by theFADEC ECU. Without pilot action, the autorelight sequence is initiated, a fuel flow rate isestablished and the ignit ion system isactivated. If a relight is achieved, FADEC willcontrol the MGT and accelerate the engineback to i ts commanded operation. TheENGINE OUT light and warning horn will turnoff after a minimum NDOT (NG accelerationspeed) or increasing MGT is established.


1-62———Rev. 2—31 JAN 2007

Figure 1-14. Engine Oil System407_MD_01_0013

SUPPLY AND

BYPASS OILPRESSURE OIL SCAVENGE OIL

SCAVENGE

RETURN

TYPICAL

VENT TO

ATMOSPHERE

TORQUE

MODULATED

PRESSURE

OIL COOLER

OIL TANK

CEFA

MAGNETIC CHIP DETECTOR

PRESSURE

REDUCER

OIL

PRESSURE SENSE

OIL FILTEROIL FILTER BYPASS VALVE

CHECK VALVE

PRESSURE REGULATING VALVE

EXTERNAL SUMP

OVERBOARD BREATHER

AIR-OIL SEPARATOR

MAGNETIC CHIP DETECTOR

TORQUEMETER PRESSURE

OIL TANK VENT

SCREEN

SCREEN


31 JAN 2007—Rev. 2———1-63

The automatic auto relight sequence willinitiate from detection of flameout until the NGspeed decays to 50%. Once the NG decaysbelow 50%, the FADEC will no longer attemptto relight the engine. In the event of anunsuccessful relight refer to BHT-407-FM-1,Section 3.

• MANUAL MODE

In MANUAL mode, the FADEC controlled autorelight circuit is disabled. Because of this, theignition system has been designed to operatecontinuously in MANUAL mode, at engine gasproducer (NG) speeds of 55% or greater toreduce the possibility of flameout.

When the FADEC system is in MANUAL mode,the NG gauge operates as a trigger device forthe engine out horn and light when NG dropsbelow 55%. In the event of a power loss to theNG indicator while operating in MANUALmode , the fa i lure mode w i l l p rov idecontinuous ignition regardless of NG speed,but the ENGINE OUT light and horn will not beactivated when NG drops below 55%.

• AUTO MODE — PILOT ASSISTED IN-FLIGHT RESTART

In addition to the above mentioned EngineRelight features, the FADEC system alsoincorporates specific relight logic, for engineout conditions, when the NG speed is between9.5% and 50%. Pilot action is required toinitiate an in-flight restart at NG speeds below50%.

When appropriate BHT-407-FM-1 proceduresare followed, the in-flight restart logic willintroduce fuel scheduling based on theexisting NG speed. Should a relight beachieved, the FADEC will accelerate theengine to an idle speed of 63% NG. As thepriority of the in-flight relight logic is to helpachieve an engine start in an emergencycondition, the hot start abort function isdisabled.

Therefore, to help reduce the possibility of anover temperature condition from occurring,

BHT-407-FM-1 procedures require that thethrottle be initially positioned to the closedposition and the start switch be positioned toSTART. Once the starter is assisting tomaintain or increase the NG speed, thethrottle can be positioned to IDLE and theFADEC will introduce fuel scheduling. Ignitionwill be provided in conjunction with activationof the starter.

As the in-flight restart logic is designed for NGspeeds between 9.5 and 50% NG, if an in-flightrestart is initiated below 9.5% NG, normal startlogic will be used to introduce fuel based on5.202 or 5.356 software (paragraph 1-12-D).Should a relight occur, the FADEC willaccelerate the engine to idle. In addition, hotstart abort logic will be enabled for startsinitiated at NG speeds below 9.5%.

1-19. AUTO RELIGHT CAUTIONLIGHT

The AUTO RELIGHT light will be ON when theignition system is activated.

1-19-A. START IN AUTO MODE

During start in the AUTO mode, the ignitionsystem is activated via the engine ignitercircuit breaker and the start circuit until theNG speed reaches 50 ±1%. Above this speed,the FADEC carries out an auto relight test andcontinues activation of the ignition systemuntil a gas producer (NG) speed of 60 ±1%, atwhich time the AUTO RELIGHT light will goOFF.

1-19-B. ENGINE OUT IN AUTO MODE

In the event of an engine out condition duringnormal engine operations with the FADEC inAUTO mode, the ignition system is activatedby the FADEC. This will occur when an engineout condition is detected and the engine isrestarted or until the gas producer (NG) speeddecays to 50%. The ENGINE OUT warninglight and horn and the AUTO RELIGHTadvisory light will be on at all times underthese conditions.


1-64———Rev. 2—31 JAN 2007

1-19-C. START AND CONTINUOUSOPERATION IN MANUAL MODE

During a start in MANUAL mode, the ignitionsystem will be activated when the starterswitch is positioned and held in START.Fol lowing the engine start and dur ingcontinuous operation at gas producer (NG)speeds above 55%, the ignition system willoperate continuously. The AUTO RELIGHTlight will be on at all times under theseconditions.

1-20. ENGINE ANTI ICE SWITCH

The ENG ANTI ICE switch, located in theoverhead console, controls the engineanti-ice bleed air solenoid valve. The engineanti-ice system will be activated when theENG ANTI ICE switch is positioned to ENGANTI ICE. This de-energizes the engineanti-ice bleed air solenoid valve allowing hotdiffuser scroll air to flow from the engineanti-icing air valve to the engine compressorfront support guide vanes and prevent theformation of ice. In the event of a totalelectrical system failure, the anti-ice systemwill fail safe to ON and provide continuousanti-icing.

When the ENG ANTI ICE switch is positionedto OFF, bus voltage is provided to the engineanti-ice bleed air solenoid valve from theengine anti-ice circuit breaker. This energizesthe engine anti-ice bleed air solenoid valveand prevents the flow of hot air from theengine anti - ice air valve to the enginecompressor front support guide vanes.

On helicopters incorporating an ENGINE ANTIICE caution panel annunciator, an engineanti-ice pressure switch is used to controlactivation of the annunciator. Engine anti-icepressure switch act ivat ion occurs onincreasing pressure at 5.5 ±.5 PSI (37.9 ±3.4kPa), which allows the ENGINE ANTI ICEannunciator to illuminate. Engine anti-icepressure switch deactivation occurs ondecreasing pressure prior to 3.0 PSI (20.68

kPa), which turns OFF the ENGINE ANTI ICEannunciator.

1-21. ENGINE INDICATORS

Engine indicators include a TORQUE, NG,MGT, NR/NP, engine oil temperature andpressure gauge, ENGINE OUT warning light,and horn, ENGINE OVSPD light, and ENGINECHIP light.

1-21-A. TORQUE GAUGE

The engine torque meter pressure sensingport outputs a specific oil pressure for aspecific engine torque. This oil pressure isfed to bo th the FADEC and a i r f rametransducers . The FADEC t ransducerinformation provides input to the ECU for theoperation of the FADEC system. The airframetransducer provides a signal to the torquegauge.

1-21-A-1. TORQUE GAUGE — RANGEMARKING

The 100% maximum on the torque gaugerepresents 560 foot-pounds (674 shafthorsepower at 100% NP) at the engine outputshaft.

This output is divided on the helicopterbetween the tail rotor drive system, the mainrotor drive system, accessory requirementsand gearbox losses.

The 93.5% on the torque gauge represents524 foot-pounds (630 shaft horsepower at100% NP). This is the maximum limit allowedon a continuous basis on the engine gearbox.The 5 minute l imit between 93.5% (524foot-pounds) and 100% (560 foot-pounds)represents a 5-minute limit on the enginegearbox in this range of power.

Any torque exceedance value recorded by theindicator represents the straight torque seenby both the engine gearbox and the maintransmission. In the event of a recordedtorque exceedance, reference to both theBHT-407-MM and Rolls Royce 250-C47B


31 JAN 2007—Rev. 2———1-65

Operat ion and Maintenance Manual isrequired to ensure appropriate inspectionsare carried out.

NOTE

Litton instrument torque does notequal FADEC ECU torque for thesame input oil pressure. This is dueto the fact that the Litton torqueinstrument and the FADEC ECU useseparate torque transducers to arriveat their respective indicated torques(i.e. engine limit = 590 foot-pounds/710 HP and transmission limit = 560foot-pounds/674 HP). To convertfrom Litton instrument% torque toFADEC ECU% torque, divide Littoninstrument% torque by 1.0535. Toconvert FADEC ECU% torque tofoot-pounds, multiply FADEC ECU%torque by 5.9. To convert fromFADEC ECU% torque to Lit toninstrument torque, multiply FADECECU% torque by 1.0535.

1-21-A-2. ENGINE TORQUE EXCEEDANCE

The torque gauge microprocessor is preprogrammed with specific torque values(Table 1-4).

There is a torque limitation of 5 minutesbetween 93.5% and 100%. To give the pilotnotice that he is approaching the end of the 5minute limit, the Trend ARC of indicator willflash and CHECK INSTR light will come onwhen 30 seconds remain in the 5 minute timeperiod.

There is a redline at 100% representing themaximum torque allowed. When the pilotexceeds 100%, the indicator Trend ARC willimmediately begin to flash and CHECK INSTRlight will illuminate.

There is an engine gearbox limit at 105.4%.When the pilot exceeds 100%, the Trend ARCof indicator will already be flashing. However,after 10 seconds at or above 105.4%, theindicator will indicate that an exceedance hasoccurred and wi l l begin to record theexceedance in NVM. At this time, the indicatorTrend ARC will stop flashing.

There is a transmission inspection limit at110%. When the pilot exceeds 100%, theTrend ARC of indicator wil l already beflashing. Any exceedance above 110% willimmediate ly be recorded in NVM. Theindicator Trend ARC will stop flashing after 10seconds above 105.4% as described above.

Table 1-4: Engine Torque Exceedance MonitoringTorque% Indication Exceedance93.6 to 100% Trend ARC begins flashing after 4.5

minutes.Not recorded.

100.1 to 105.3% Trend ARC begins flashing immediately and flashes continuously as long as in this range.

Not recorded.

105.4 to 110% Trend ARC begins flashing immediately. Flashes for 10 seconds, then stops flashing.

Recorded after 10 seconds.

Above 110.1% Trend ARC flashes prior to reaching 110.1%, as described above, but stops flashing at 110.1% or above.

Recorded immediately.


1-66———Rev. 2—31 JAN 2007

1-21-B. GAS PRODUCER GAUGE (NG)

The gas producer (NG) gauge displays enginegas producer speed in percent of rated RPM(Table 1-5).

An NG Speed Pickup mounted in the enginegearbox provides two separate signals to theECU. One signal is the primary driver of theNG gauge and is also a secondary signal forthe ECU. The second signal is the primaryinput to the ECU. The primary signal to the

gauge will be passed through the ECU, evenwhen the ECU is not powered.

When the FADEC system is in MANUAL mode,the NG gauge operates as a trigger device forthe ENGINE OUT horn and light when NGdrops below 55%. In the event of a power lossto the NG indicator whi le operat ing inMANUAL mode, the failure mode will providecontinuous ignition regardless of NG speed,but the ENGINE OUT light and horn will not beactivated when NG drops below 55%.

1-21-C. DUAL TACH GAUGE

The Dual Tachometer has two pointers: onethat displays Main Rotor RPM on the outerscale and one that displays engine NP on theinner scale. All scales are in percent of ratedRPM. If Quiet Cruise Mode kit is installed,refer to the BHT-407-FMS-25.

The NP side of the instrument is powered bythe 28 VDC bus through its own circuitbreaker.

An NP Speed Pickup mounted on the engineprovides two separate signals to the ECU.One signal is the primary driver of the NPgauge and also a secondary signal for theECU. The second signal is the primary inputto the ECU. The primary signal to the gaugewill be passed through the ECU, even whenthe ECU is not powered.

The NR side of the gauge is powered by the 28VDC bus through its own circuit breaker.

An N R speed p ickup mounted on thetransmission lower case provides threeseparate identical signal outputs of rotorRPM. One signal is sent to the NR circuit ofthe dual tachometer gauge. One signal is sentto the ECU and the other is sent to the MainRotor RPM sensor switch.

1-21-D. MEASURED GAS TEMPERATUREGAUGE

The measured gas temperature (MGT) gaugedisplays engine gas temperature of airbetween gas producer turbine and powerturbine in degrees Celsius.

The MGT gauge has one set of exceedancelevels for starting and another for normaloperation (Table 1-6 through Table 1-9). Theindicator uses the start exceedance levelswhen the FADEC/START relay is engagedduring normal starts.

Table 1-5: GAS PRODUCER GAUGE (NG) Exceedance Monitoring

NG% Indication Exceedance

105.1 to 106% Trend ARC starts flashing immediately. Stops flashing after 10 seconds.


Above 106.1% Trend ARC begins flashing prior to reaching 106.1%, as described above. Stops flashing at 106.1% and above.



31 JAN 2007—Rev. 2———1-67

Table 1-6: MGT Exceedance Monitoring — During Start (P/N 407-375-001-101/103)MGT°C Indication Exceedance826 to 926°C Trend ARC begins flashing after 6 seconds.

Stops flashing after 10 seconds. Recorded after 10 seconds.

926.1 to 927.9°C Trend ARC begins flashing prior to reaching 926.1°C, as described above but stops flashing at 927.9°C and above.

Recorded after 1 second.

Above 928°C Trend ARC begins flashing prior to reaching 928°C, as described above but stops flashing at 928°C or above.


Table 1-7: MGT Exceedance Monitoring — During Normal Operation (P/N 407-375-001-101/103)MGT°C Indication Exceedance727.1 to 779°C Trend ARC begins flashing after 4.5

minutes. Stops flashing after 5 minutes.Recorded after 5 minutes.

779.1 to 826°C Trend ARC begins flashing immediately and stops flashing after 12 seconds.


Above 826.1°C Trend ARC begins flashing prior to reaching 826.1°C, as described above but stops flashing at 826.1°C or above.


Table 1-8: MGT Exceedance Monitoring — During Start (P/N 407-375-001-105 and Subsequent)MGT°C Indication Exceedance843 to 926°C Trend ARC begins flashing after 6 seconds.

Stops flashing after 10 seconds.Recorded after 10 seconds.

926.1 to 927.9°C Trend ARC begins flashing prior to reaching 926.1°C, as described above, but stops flashing at 927.9°C and above.

Recorded after 1 second.

Above 928°C Trend ARC begins flashing prior to reaching 928°C as described above, but stops flashing at 928°C or above.


Table 1-9: MGT Exceedance Monitoring — During Normal Operation(P/N 407-375-001-105 and Subsequent)

MGT°C Indication Exceedance727.1 to 779°C Trend ARC begins flashing after 4.5

minutes. Stops flashing after 5 minutes.Recorded after 5 minutes.

779.1 to 905°C Trend ARC begins flashing immediately and stops flashing after 12 seconds.



1-68———Rev. 2—31 JAN 2007

1-21-E. ENGINE OIL TEMPERATURE/PRESSURE GAUGE

The engine oil temperature and pressuregauge is a dua l inst rument thatsimultaneously displays oil temperature indegrees Celsius on the right side display andoil pressure in PSI on the left side display.Each side of the indicator is powered by itsown c i rcu i t breaker. The eng ine o i ltemperature input signal is provided by athermobulb installed on the engine oil tank.The engine oil pressure is provided by atransducer mounted on the forward enginefire wall.

1-22. ENGINE OUT WARNING LIGHTAND HORN

The ENGINE OUT light and (pulsing) warninghorn circuit will activate when the FADECdetects an engine flameout (by sensing NGdeceleration) or when gas producer (NG)speed is 55 ±1% or less.

1-22-A. FADEC IN AUTO MODE

If the FADEC detects an engine flameout (NGdeceleration) or NG speed of 55 ±1% or less,the ENGINE OUT light and (pulsing) warninghorn circuit will activate. In addition, theFADEC will automatically initiate an autorelight sequence immediately upon detectionof a flameout (NG deceleration).

1-22-B. FADEC IN MANUAL MODE

In the MANUAL mode, the FADEC engine outdetection circuit is deactivated and activationof the ENGINE OUT light and (pulsing)warning horn is control led by the gasproducer NG gauge (NG less than 55 ±1%).

Engine out (flameout) protection in theMANUAL mode is provided by continuousactivation of the ignition system at NG speedsof 55% or greater.

The engine out (pulsing) warning horn can bemuted in either the AUTO or MANUAL modeby pressing the HORN MUTE switch.

1-23. ENGINE OVERSPEEDWARNING LIGHT

The ENGINE OVSPD light will illuminate if theFADEC detects a NG overspeed of 110 ±1% ora NP overspeed of 118.5 ±1% (5.202 and 5.356FADEC software). Illumination occurs whenthe overspeed solenoid valve is activatedwithin the Hydro Mechanical Unit (HMU). TheENGINE OVSPD light will also illuminate whenNP versus TORQUE (5.202 FADEC software) isabove the maximum continuous limit (102.1%NP at 100% torque to 108.6% NP at 0% torque)or when NP is above the maximum continuouslimit of 102.1% (5.356 FADEC software). TheENGINE OVSPD light is operational with theFADEC in AUTO or MANUAL mode.

If the ENGINE OVSPD light is activated duringengine operation, due to an exceedance, itwill be recorded by the ECU and the pilot willbe provided with a maintenance advisory onshutdown in the form of a FADEC DEGRADEDlight. The FADEC DEGRADED light wil lilluminate when NG speed decays below 9.5%.If the pilot fails to recognize illumination ofthe FADEC DEGRADED light on shutdown, itwill be illuminated the next time electricalpower is applied following the FADEC systemself test. When the FADEC DEGRADED light isil luminated as a maintenance advisory,maintenance action is required prior to furtherflight. Peak values of exceedances are located

Above 905°C Trend ARC begins flashing prior to reaching 905°C, as described above, but stops flashing at 905°C or above.


Table 1-9: MGT Exceedance Monitoring — During Normal Operation(P/N 407-375-001-105 and Subsequent) (Cont)

MGT°C Indication Exceedance


30 APR 2008—Rev. 4———1-69

on the Engine History Data page of theEMC-35A Maintenance Terminal.

The ENGINE OVSPD light will also illuminatemomentarily during the overspeed shut downtest when the overspeed solenoid valve isactivated within the HMU (BHT-407-FM-1).

1-24. ENGINE CHIP CAUTION LIGHT

The ENGINE CHIP light will il luminate ifmetallic particles in the oil accumulate oneither the engine sump or scavenge chipdetectors. The engine sump chip detector islocated on the lower left side of the enginegearbox and the engine scavenge chipdetector is located on the forward right-handside of the engine gearbox. Both chipdetectors are a quick disconnect design.

1-25. FUEL SYSTEM DESCRIPTION

The fuel system (Figure 1-15) consists of twocrash resistant, bladder type fuel cells. Theforward fuel cell is located underneath andbetween the aft facing passenger seats. Theaft fuel cell is located underneath and behindthe aft passenger seats.

Both fuel cells are serviced through the fillerport located on the right side of the helicopter.Approximately 28.4 US gallons (193.1 lb) offuel will accumulate in the aft main fuel cell,prior to the forward cell being filled throughthe gravity feed stand pipe. The gravity feedstand pipe connects the aft fuel cell to theforward tank. As the aft fuel cell fills beyondthe level of the top of the stand pipe, theforward tank will be completely filled (forwardtank fuel capacity 37.6 US gallons (256.0 lb).The aft tank will then be filled to the level ofthe filler port (usable fuel capacity 127.8 USgallons (869.0 lb).

If the auxiliary tank is installed, it will be filledat the same time as the aft tank through twoopenings located on the lower aft wall of the

main fuel cell, which are connected to thebot tom of the aux i l ia ry tank ( re fer toBHT-407-FMS-6 for information on this kit).

Fuel from the forward tank is transferred tothe main tank by two transfer pumps mountedon a sump plate assembly located on thebottom of the forward fuel cell. Fuel from themain fuel cell will be supplied to the enginethrough two boost pumps located at the baseof the main fuel cell on a sump plate. The fuelfrom the two boost pumps joins into acommon fuel line that passes through a fuelshutoff valve, then through an airframemounted fuel filter before reaching the enginedriven pump on the HMU. A solenoid sumpdrain valve is installed on the sump plateassemblies of both the forward and aft fueltanks. The drain valves are activated by twoswitches located on the right hand side of thelower aft fuselage. These switches aredeactivated when the fuel valve switch is ONto prevent inadvertent activation during flight.

1-25-A. FUEL SYSTEM OPERATION

With power applied to the helicopter and theFuel XFR/Boost Right and Fuel XFR/BoostLeft circuit breaker switches ON, the twotransfer pumps and the two boost pumps areoperating. Refer to Figure 1-16 for the fueltransfer system schematic.

Each transfer pump sends fuel through aone-way check valve located at the outlet ofeach pump and a common tee fitting througha line which transfers the fuel into the aft tank.Each check valve will assure that if either ofthe transfer pumps becomes inoperative fuelwill be pumped to the aft tank and not throughthe inoperative pump back into the fwd fueltank.

A modified tee fitting incorporating an orificeallows for a specific amount of fuel to bleedback into the forward tank to reduce thetransfer rate.


1-70———Rev. 2—31 JAN 2007

Figure 1-15. Fuel System407_MD_01_0014

VENT LINE

QTY PROBE(3 PLACES)

TRANSFERBLEED

TRANSFERLINE

CHECK VALVEWITH THERMALRELIEF

TRANSFERPUMP DUAL

ELECTRICSUMP DRAIN

PRESSURESWITCH

LOW LEVELSWITCH

FUEL VALVE SWITCH

FIREWALL

AIRFRAMEFUEL FILTER

MANUALDRAIN VALVE

ENGINE

FUEL VALVE(AMBER)

FUEL SIGNAL

CONDITIONER

R/FUEL XFR(AMBER)

L/FUEL XFR(AMBER)

L/FUEL BOOST

(AMBER)FUEL LOW(AMBER)

R/FUEL BOOST

(AMBER)

FUEL

PRESS

FUELQTYLBS

FUEL FILTER

(AMBER)

DRAINHOSE

AUX FUEL TANKKIT (REF)

OVERBOARDVENT

CHECK VALVEWITH THERMALRELIEF

BOOST PUMP DUAL

LOW LEVEL SWITCH

ELECTRIC SUMPDRAIN

TYPICAL

INTERCONNECT LINE

SHUT OFF VALVE WITHTHERMAL RELIEF

PRESSURETRANSDUCER

FILLER CAP

ENGINE FEEDLINE

AFT FUELCELL

FWD FUELCELL


31 JAN 2007—Rev. 2———1-71

Figure 1-16. Fuel Transfer System Schematic (Sheet 1 of 2)407_MD_01_0015

TYPICAL

ON

OFF4

5

6

28 VDC BUS

FUEL XFR/

BOOST RIGHT

FUEL XFR/

BOOST LEFT

LOW LEVEL

DETECTOR

FUEL SIGNAL

CONDITIONER

POWER INTIME

DELAY

TIME

DELAYPOWER IN

FORWARD FUEL TANK

FUEL

PROBE

FLOAT

RIGHT XFR

PUMP RELAY

LEFT XFR

PUMP RELAY

BATTERY SWITCH

BATTERY POWER

28 VDC POWER

S/N 53000 THROUGH 53174

RIGHT FUEL

PRESSURE SWITCH

LEFT FUEL

PRESSURE SWITCH

RIGHT XFR

PUMP

LEFT XFR

PUMP

CAUTION

WARNING PANEL

R/FUEL XFR

L/FUEL XFRD

L

PUMP OUTPUT

PRESSURE

PUMP OUTPUT

PRESSURE

FUEL QTY IND

3

10

10


1-72———Rev. 2—31 JAN 2007

Figure 1-16. Fuel Transfer System Schematic (Sheet 2 of 2)407_MD_01_0016

TYPICAL

ON

OFF4

5

6

28 VDC BUS

FUEL XFR/

BOOST RIGHT

FUEL XFR/

BOOST LEFT

LOW LEVEL

DETECTOR

FUEL SIGNAL

CONDITIONER

POWER INTIME

DELAY

TIME

DELAY

GROUND CIRCUIT

FOR XFR LIGHTS

BASED ON

ACTIVATION TABLE

POWER IN

FORWARD FUEL TANK

FUEL

PROBE

FLOAT

RIGHT XFR

PUMP RELAY

LEFT XFR

PUMP RELAY

BATTERY SWITCH

BATTERY POWER

28 VDC POWER

S/N 53175 AND SUBSEQUENT

RIGHT FUEL

PRESSURE SWITCH

LEFT FUEL

PRESSURE SWITCH

RIGHT XFR

PUMP

LEFT XFR

PUMP

CAUTION

WARNING PANEL

R/FUEL XFR

L/FUEL XFRD

L

PUMP OUTPUT

PRESSURE

PUMP OUTPUT

PRESSURE

FUEL QTY IND

3

10

10


31 JAN 2007—Rev. 2———1-73

Each boost pump sends fuel through aone-way check valve located at the outlet ofeach pump and a common tee fitting. It thenflows through a common line to a fittinglocated at the top right side of the main fuelcell. The line leaves the fuel cell and passesthrough the fuel shut off valve. A pressuretransducer (located between the main fuel cellfitting and the fuel shut off valve) supplies theelectrical signal to the instrument panelmounted fuel pressure gauge.

During normal operation, fuel will first beused from the aft fuel cell until its level isequivalent to the top of the forward tank. Atthis time, the aft cell and forward cell will beused equally down to the level of the top ofthe gravity feed stand pipe. At this point, allfuel will be used from the forward tank.Finally, after the forward tank is empty, theremaining fuel in the aft tank will be used.

Refer to paragraph 1-25-P-2 for operationalinformation of the L/FUEL XFR and R/FUELXFR LIGHTS and paragraph 1-25-R forinformation on the FUEL LOW LIGHT.

1-25-B. LEFT FUEL BOOST/XFRALTERNATE ELECTRICALCIRCUIT

In the event that a short circuit or battery hotcondition occurs in the helicopter and all DCbus power is shut off, it is desirable tomaintain the operation of one transfer pumpand one boost pump.

The BATT switch configures the DC powerfeed to the left fuel boost pump and the leftfuel transfer pump between the DC bus andthe helicopter battery (Figure 1-17).

In the event the battery switch is positioned toOFF dur ing he l icopter operat ions , analternate circuit is provided to allow operationof the left fuel transfer and left fuel boostpumps. During this condition, with the fuelvalve switch positioned to ON, battery voltage

is supplied through the FUEL BOOST/XFRbackup circuit breaker, the fuel valve switch,the battery switch, and the left fuel XFR/boostcircuit breaker switch to the left fuel transferand left fuel boost pumps.

1-25-C. FUEL SYSTEM CONTROLS

Fuel system controls are located on theinstrument panel, overhead console, and aftlower right fuselage. The controls consist of aFUEL VALVE switch, a FUEL BOOST/XFRLEFT switch, a FUEL BOOST/XFR RIGHTswitch, and two FUEL CELL DRAIN switches.

1-25-D. FUEL VALVE SWITCH

The FUEL VALVE switch is located on thelower right side of the instrument panel. It is aguarded two position toggle switch thatprovides a means of shutting off the flow offuel to the engine.

The fuel shutoff valve is a motorized gatevalve which must be driven to either the openor closed position. When activated by theFUEL VALVE switch, the shutoff valve motorwill rotate in the appropriate direction toOPEN or CLOSE the gate valve. The FUELVALVE light will momentarily illuminateduring valve transit.

1-25-E. LEFT AND RIGHT FUEL BOOST/XFR SWITCH

The left and right FUEL BOOST/XFR switches(Figure 1-16) are a circuit breaker toggledesign and located on the overhead console.Each switch controls the operation of itsrespective boost pump located in the mainfuel tank and transfer pump in the forwardfuel tank. In addition to the left and rightFUEL/XFR switches, separate electricalcircuits within the fuel signal conditioner areused to control the operation of the left andright fuel transfer pumps as the forward fueltank is emptied.


1-74———Rev. 2—31 JAN 2007

The LEFT FUEL/XFR switch is outlined by ayellow border (Figure 1-5) to identify that ithas an alternate circuit (Figure 1-17). In theevent the BATT switch is positioned to OFFduring helicopter operations, an alternatecircuit is also provided to allow operation ofthe left fuel transfer and left fuel boost pumps.During this condition, with the FUEL VALVEswitch positioned to ON, battery voltage issupplied through the FUEL BOOST/XFRBACKUP circuit breaker (located on the leftside of the instrument pedestal above the chinbubble), the FUEL VALVE switch, the BATTSWITCH, and the LEFT FUEL BOOST/XFRcircuit breaker switch to the left fuel transferand left fuel boost pumps. Therefore, in thissituation the LEFT BOOST pump will continueto run. The LEFT XFR pump will continue torun as long as the fuel signal conditionerdetermines that there is fuel in the forwardtank.

1-25-F. FUEL CELL DRAIN SWITCHES

The FUEL CELL DRAIN switches provide ameans of draining fuel from both the forwardand aft fuel cells individually. The switchesare a momentary push button design made upof an outer environmental seal and an innerswitch assembly. They are mounted side byside on the lower right side of the aft fuselage,below and aft of the fuel filler port.

The FUEL VALVE switch must be in the OFFposition to operate either of the fuel cell drainvalves. This prevents inadvertent operation ofthe fuel cell drains during flight.

With helicopter electrical power provided andthe FUEL VALVE switch positioned to OFF,pressing the FUEL CELL DRAIN switches willOPEN the respective drain valves allowingfuel to be drained from the forward and mainfuel cells.

1-25-G. FUEL SYSTEM INDICATORS

Fuel system indicators consist of a fuelquantity indicator, fuel pressure indicator,FUEL VALVE light, L/FUEL BOOST light, R/FUEL BOOST light, L/FUEL XFR light, R/FUEL XFR light, FUEL FILTER light, and aFUEL LOW light.

1-25-H. FUEL SYSTEM CAPACITY

For fue l system capac i ty re fer to theBHT-407-FM-1, Section 5.

1-25-I. FUEL QUANTITY GAUGINGSYSTEM (FQGS)

The fuel quantity gauging system (FQGS)measures the quantity of fuel in the two mainfuel tanks. The FQGS also measures thequantity of fuel in the auxiliary fuel tank whenit is installed.

The fuel quantity is measured by threecapacitance type probes in the fuel tanks. Thesignals from the probes are used by the fuelsignal conditioner to calculate the fuel weight.The signal conditioner provides a signal tothe fuel quanti ty gauge to display thecalculated fuel weight.

1-25-J. FUEL QUANTITY SIGNALCONDITIONER

The signal conditioner is a separate unit thatis located on the aft electrical equipment shelfnext to the DC controller (Voltage Regulator).

The electronic interface circuits for the fuellow level detection system are located alongwith the fue l gaug ing system s igna lconditioner circuits in the same physical unit.

Both systems are physically and electricallyseparate within the unit. There are three LEDslocated on the a ft s ide o f the s igna lconditioner which are used to display theFQGS status for purposes of troubleshooting.


31 JAN 2007—Rev. 2———1-75

Figure 1-17. Left Fuel Boost/XFR Alternate Circuit Schematic407_MD_01_0017

TYPICAL

ON

ON

1

2

3

4

5

6 4

5

6

OFF

OFF

15

10

BATTERY

BATTERY RELAYFUEL BOOST/

XFR BACKUP

FUEL XFR/

BOOST LEFT

LEFT

BOOST PUMP

LEFT

TRANSFER PUMP

LEFT XFR

PUMP RELAY

FUEL SIGNAL

CONDITIONER

BATTERY SWITCH

28 VDC BUS

FUEL VALVE SWITCH

A2

- +


1-76———Rev. 2—31 JAN 2007

1-25-K. SIGNAL CONDITIONERBUILT-IN-TEST (BIT)

1-25-K-1. POWER-UP BIT

The signal conditioner receives power fromthe 28 VDC bus through the FUEL QTY INSTRcircuit breaker. The signal conditioner carriesout a power up BIT when the unit is firstprovided power. The power-up BIT must becompleted before the signal conditioner cantake any readings of fuel quantity. The signalconditioner should complete a power-up BITcheck within approximately 4 seconds afterapplication of power. There is no connectionbetween the BIT feature o f the s ignalconditioner and the BIT performed by the fuelquantity gauge.

If a failure is detected during the power-up BITor if an error is found in the probe signalreceived or the power source for the probeinput, the signal conditioner will blank the fuelquantity gauge.

If errors have been detected and the gaugehas been blanked, the signal conditioner willnot turn the gauge back on even if the errorhas been corrected unless the power isturned OFF and ON to the signal conditioner.

In addition, the failures detected will bedisplayed on three LEDs on the back of thesignal conditioner (refer to the BHT-407-MMfor fault explanation).

1-25-K-2. CONTINUOUS BIT

The s igna l cond i t ioner ca r r ies ou t acontinuous BIT whenever it is powered. If afailure is detected during the continuous BITor if an error is found in the probe signalreceived or the power source for the probeinput, the signal conditioner will blank thegauge display.

If errors have been detected and the gaugehas been blanked, the signal conditioner willnot turn the gauge back on even if the error

has been corrected unless the power isturned OFF and ON to the signal conditioner.

In addition, the failures detected will bedisplayed on three LEDs on the back of thesignal conditioner (refer to the BHT-407-MMfor fault explanation).

1-25-L. FUEL QUANTITY CALCULATION

A microprocessor in the signal conditioneruses the information provided by the threefuel probes to compute the weight of the fuelin the following steps.

1. Uses the Main Tank Forward FuelProbe (Probe No. 2) input signal fordensity correction if it is totallyimmersed in fuel or if not, uses adefault density (default density is6.6594 pounds/gallons).

2. Calculates the height of the fuelindicated for each of the three probescorrected for fuel density by usingthe value from step 1.

3. The calculated height on each probeis used to look up a volume of fuel ingallons in a table contained in theNVM of the signal conditioner.

4. The weight of the fuel is thencalculated by multiplying the volumeby the density. A calculation is donefor all three probes to compute thetotal system weight of fuel. Acalculation is also done for only theForward Tank Fuel Probe (ProbeNo. 3).

To ta l system weight in format ion istransmitted from the signal conditioner to thefuel quantity gauge. When the forward fuelquantity switch is pushed, the weight for theforward tank only is transmitted.

1-25-M. FUEL QUANTITY GAUGE

The usable fue l weight ( in pounds) iscalculated by the fuel signal conditioner andis displayed by the gauge.


30 APR 2008—Rev. 4———1-77

1-25-N. FUEL QUANTITY BUTTON

The fuel quantity gauge normally indicatesthe total usable fuel in both fuel tanks andauxiliary tank (if installed). Pushing the FUELQTY FWD TANK button will make the fuelquantity gauge display the fuel in the forwardtank only.

1-25-O. FUEL PRESSURE/AMMETERGAUGE

The fuel pressure/ammeter gauge is a dualdisplay gauge. The left side of the instrumentdisplays the pressure output from the two fuelboost pumps in pounds per square inch (PSI).The fuel pressure side of the instrument ispowered by its own circuit breaker. Theinstrument receives its input signal from atransducer mounted between the fuel boostpumps and the fuel shut off valve. Theminimum pressure limit is set to ensure thatsufficient fuel pressure will be supplied to theinput of the engine driven fuel pump.

1-25-P. FUEL VALVE LIGHT

The FUEL VALVE light will illuminate when thefuel shut off valve is in transit or has stoppedsomewhere between the full OPEN or fullCLOSED position.

1-25-P-1. L/FUEL AND R/FUEL BOOSTLIGHTS

The L/FUEL BOOST and R/FUEL BOOSTlights will illuminate when their respectivefuel pressure switch senses a decreasingboost pump output pressure of 1.5 ±0.5 PSI.When boost pump pressure is increasing, thelights will extinguish prior to the pressurepassing through 5 PSI.

1-25-P-2. L/FUEL XFR AND R/FUEL XFRLIGHTS

1-25-P-2-A. S/N 53000 THROUGH 53174

The L/FUEL XFR and R/FUEL XFR lights(Figure 1-16, Sheet 1) will illuminate when

their respective fuel pressure switch senses adecreasing boost pump output pressure of1.5 ±0.5 PSI. When boost pump pressure isincreasing the lights will extinguish prior tothe pressure passing through 5 PSI.

Fuel must be present in the forward fuel cellfor the annunciator circuits to operate sincethey are both controlled by the fuel signalconditioner. When the forward fuel cell isnearing depletion the signal conditioner usestime delays to control the transfer pumps andlights. The time delay allows continuedoperation of the transfer pumps to ensure allof the fuel in the forward tank is transferred tothe main fuel tank. The forward fuel cell will beempty when approximately 193.1 pounds oftotal fuel is indicated.

The fuel signal conditioner time delay whichcontrols the L/FUEL XFR pump circuit isprovided by the FUEL XFR/BOOST LEFTcircuit breaker. The fuel signal conditionertime delay which controls the R/FUEL XFRpump circuit is provided by the FUEL XFR/BOOST RIGHT circuit breaker.

The L/FUEL XFR and R/FUEL XFR lightcircuits will remain operational during thetime delay and will illuminate in the eventtransfer pump output pressure drops below1.5 ±0.5 PSI.

Once the time delay periods are ended, thetransfer pumps and the L/FUEL XFR and R/FUEL XFR light circuits are deactivated. Thetransfer pump and light circuits will stayinoperative as long as the forward fuel cell isempty. Reactivation of the transfer pumpcircuits will occur when approximately 18pounds of fuel enters the forward fuel cell. Asthe forward fuel cell is empty at approximately193.1 pounds of fuel, if the helicopter is shutdown at, or being refueled to between 193.1and approximately 211.1 pounds of total fuel,it is possible that up to 18 pounds of fuel mayremain in the forward fuel cell as unusable.


1-78———Rev. 4—30 APR 2008

1-25-P-2-B. S/N 53175 AND SUBSEQUENT

Activation of either annunciator circuit isdependent on the relationship between thefuel quantity in the forward fuel tank and thetotal quantity of the fuel system.

The L/FUEL XFR or R/FUEL XFR transfer light(Figure 1-16, Sheet 2) will illuminate whentheir respective fuel pressure switch senses adecreasing boost pump pressure of 1.5 ±0.5PSI, provided the condit ions shown inTable 1-10 (Transfer Light Activation Table)are met:

With a L/FUEL XFR or R/FUEL XFR pumppressure of 5 PSI or greater, the respectiveleft and right fuel pressure switches will openand cause the L/FUEL XFR or R/FUEL XFRlights to go off.

When the fuel signal conditioner detects aforward fuel tank quantity of less than 25pounds and a total fuel system quantity ofless than 250 pounds, for 10 consecutiveseconds, L/FUEL XFR or R/FUEL XFR transferl ights wi l l be deac t ivated to p reventin termit tent l ight f l icker ing whi le theremaining fuel is transferred from the forwardfuel tank to the aft fuel tank.

When forward fuel tank depletion is detectedby the number 3 fuel probe and the low leveldetector, the input signals to the fuel signalconditioner are removed. With the inputsignals removed, the fuel signal conditionerutilizes two 360 second time delays prior toremoving the ground to the right and lefttransfer pump relays. This allows the rightand left transfer pumps to continue runningfor 360 seconds to ensure all the fuel in the

forward tank is transferred to the main fueltank. The forward fuel cell will be empty whenapproximately 193.1 pounds of total fuel isindicated.

The annunciator and transfer pump circuitswill stay inoperative until the fuel system isrefueled with an appropriate amount of fuel toreactivate the system. Reactivation of thetransfer pump circuits will occur whenapproximately 18 pounds of fuel enters theforward fuel cell. As the forward fuel cell isempty at approximately 193.1 pounds of fuel,if the helicopter is shut down at, or beingrefueled to between 193.1 and approximately211.1 pounds of total fuel, it is possible thatup to 18 pounds of fuel may remain in theforward fuel cell as unusable.

1-25-Q. FUEL FILTER LIGHT

The FUEL FILTER light will illuminate if theairframe fuel filter is in an impending bypasscondition.

This will occur when a differential pressure of0.875 ±0.125 PSI is present between the input

Table 1-10: Transfer Light Activation Table, S/N 53175 and SubsequentFWD TANK< 25 POUNDS FOR 10 CONSECUTIVE SECONDS

TOTAL FUEL< 250 POUNDS FOR 10 CONSECUTIVE SECONDS

L/FUEL XFR OR R/FUEL XFR LIGHT ILLUMINATED (AT A DECREASING PRESSURE OF 1.5 ±0.5 PSI)

FALSE FALSE YESFALSE TRUE YES TRUE FALSE YESTRUE TRUE NO


31 JAN 2007—Rev. 2———1-79

and output of the fuel filter. The airframe fuelfilter will go into bypass at 3.75 ±0.25 PSI.

1-25-R. FUEL LOW LIGHT

The FUEL LOW light circuit is designed toalert the pilot that the main fuel tank quantityis low. I l luminat ion w i l l occur whenapproximately 100 ±10 pounds of usable fuelremains in the main fuel cell.

A float type low level switch is used to detectthe fuel low condition. The input from the lowlevel detector is passed through the fuelsignal conditioner which provides a 13 ±3second time delay to reduce the possibility ofintermittent annunciator flickering due to fuelsloshing.

In addition, momentary activation (less than 2seconds) of the FUEL LOW light may bev is ib le when power is app l ied to thehelicopter and fuel quantity is greater than therequired low level activation point. This is anormal occurrence and is a function of fuelsignal conditioner power up logic.

1-26. RETIREMENT INDEX NUMBER(RIN)

Each component with a ret irement l i fesensitive to "TORQUE EVENTS" will beassigned a maximum RETIREMENT INDEXNUMBER (RIN). This RIN corresponds to themaximum allowed fatigue damage resultingfrom lifts and takeoffs. A new component willbegin with an accumulated RIN of zero thatwill be increased as lifts and takeoffs areperformed. The operator will record thenumber of lifts and takeoffs and increase theaccumulated RIN accordingly. When themaximum RIN is reached, the component willbe removed f rom serv ice . Cer ta incomponents may be assigned a life in hoursin addition to the RIN.

Pilots are to record “TORQUE EVENTS” foreach flight. Maintenance personnel will

conver t the “TORQUE EVENTS” in to“RETIREMENT INDEX NUMBERS” (RIN) totrack the lives of all required components.

A "TORQUE EVENT" is defined as a takeoff(one takeoff plus the subsequent landing =one RIN) or a lift (internal or external). Forexample, if an operator performs six takeoffsand ten sling loads, this would total 16 torqueevents: (6 takeoffs = 6 events, 10 sling loads =10 events, 6 + 10 = 16 events total).

1-27. TRANSMISSION

The t ransmiss ion and mast assembly(Figure 1-18) transfers the engine torque tothe main rotor system with a two stage gearreduction of 15.29 to 1.0 (6317 to 413 RPM).

The transmission assembly is made up of atop support case and lower case whichcontains an input pinion and bevel geararrangement, a planetary gear train, and anaccessory gear drive. The components thatare attached to the transmission and mastassembly are the engine to transmissiondr iveshaft , t ransmiss ion o i l pump,transmission oil filter housing, hydraulicpump, rotor RPM monopole pickup, and twoelectric chip detectors.

The transmission assembly is attached to theroof of the helicopter, forward of the engineby a pylon installation. The pylon installationuses two side beams, four elastomeric cornermounts, and two for/aft restraint springs.

1-27-A. FREEWHEEL ASSEMBLY

The freewheel unit (Figure 1-18) is mountedon the engine gearbox, and is driven underpower from the engine power takeoff gearshaft. Engine power is transmitted to theouter race of the freewheel unit, then throughthe engaged sprag clutch which drives thefreewheel inner shaft and couples the engineto the transmission input drive shaft.


1-80———Rev. 2—31 JAN 2007

Figure 1-18. Transmission Assembly407_MD_01_0018

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

11.

12.

13.

Filler cap

Corner mount

Mast assembly

Monopole sensor (NR)

Pylon beam assembly

Upper chip detector

Transmission assembly

Engine-to-transmission driveshaft

Disc (standard)

Freewheel assembly

Freewheel chip detector

Rotor brake kit (optional)

Lower chip detector

3

7

8

12

9

10

6

13

11

4

5

2

1

31 JAN 2007 Rev. 2 1-81


The tail rotor drive system is driven through aflexible coupling and a splined adaptermounted on the aft end of the freewheel innershaft.

Dur ing autorota t ion , the sprag c lu tchdisengages and the rotational forces of themain ro tor are a l lowed to d r ive thetransmission mounted accessories, and tailrotor drive system.

1-27-B. TRANSMISSION OIL SYSTEM

The transmission oil system (Figure 1-19)lubricates the transmission, mast, andfreewheel assemblies. An oil level sight gaugeis located on the right side of the transmissionlower case and may be viewed through acutout in the air induction cowling. A non-vented filler cap is located on the right side ofthe top case to fill the transmission oil system.

The transmission accessory gear drives theoil pump which delivers 6.0 to 6.7 GPM. Thepressure is controlled by a pressure regulatorvalve which is set at approximately 52 PSI. Thepump scavenges oil from the lower case sumpthrough a wire screen and the lower chipdetector. It is then directed to the transmissionmounted oil manifold and the filter element.

To ensure oil flow is not restricted, the filterincorporates an impending bypass indicatorbutton on the end of the filter housing whichwill extend at 14 ±2 PSI. Ensure appropriatemaintenance actions are carried out followingan impending bypass indication in accordancewith the BHT-407-MM. Additionally, a filterbypass valve will open if the differentialpressure reaches 17 PSI and if the differentialpressure reaches 29.6 to 38.6 PSI, a highpressure relief valve will open and the oil willtotally bypass the filter.

The transmission mounted oil manifold alsoincorporates a thermostatic valve which

controls the flow of oil to the oil cooler. At oiltemperatures below 150°F (66°C), the oilcooler is bypassed and the oil is directed backto the transmission. As the oil temperatureincreases above 150°F (66°C) oil is graduallydirected to the oil cooler until at 180°F (82°C),all oil is directed through the cooler.

A temperature bulb for the oil temperatureindicator and a thermoswitch for the XMSNOIL TEMP l ight are also located on thetransmission mounted oil manifold.

After the oil exits the cooler (or bypasses thecooler through the thermostatic bypass valve),it is then directed to the transmission tolubricate the various gears and bearings. Theoil is then directed to a deck mounted oilmanifold located below the transmission inputdriveshaft. A pressure transducer and oilpressure swi tch are mounted on thet ransmiss ion deck o i l mani fo ld . Thetransducer provides signals to the oil pressuregauge and the pressure switch controls theXMSN OIL PRESS light.

After leaving the deck mounted oil manifold,the oil flows through the forward fire wall anda tee fitting equipped with two restrictors. Therestrictor fittings reduce the flow and directthe oil to the freewheel forward duplex bearingand aft housing bearing. The oil that lubricatesthe bearing in the aft housing moves forwardthrough the hollow engine output driveshaft tothe freewheel sprag clutch and bearing. Theoil is then collected in the forward freewheelhousing and re tu rned to the maintransmission lower case at atmosphericpressure.

1-27-C. TRANSMISSION INDICATORS

Transmission indicators include an oi ltemperature and pressure gauge, XMSN OILTEMP light, XMSN OIL PRESS light, and XMSNCHIP light.

1-82 Rev. 2 31 JAN 2007


Figure 1-19. Transmission Oil System407_MD_01_0019

CHIP DETECTOR AND

OIL INLET SCREEN

Supply oil at pump pressure, 80 to 150 PSI.

Supply oil at regulated delivery pressure, 50 to 55 PSI.

Return oil at atmospheric pressure.

Scavenge oil

LEGEND

OIL LEVEL

SIGHT GAUGE

THERMOSWITCH

ACTUATES AT 230°F

TEMP BULB

2HIGH PRESSURE

RELIEF VALVE

(CRACKS AT 29.6

TO 38.6 PSI)

OIL

COOLER

OIL FILTER

OIL SUMP

PRESSURE

REGULATOR

(REGULATES OIL

PRESSURE FROM

50 TO 55 PSI)

THERMOSTATIC

BY-PASS VALVE

(CRACKS AT 190°F

MIN. AT 40 PSID

MIN.) (STARTS

CLOSING AT 150°F

AND IS CLOSED

AT 178°F ± 2°F

OIL SUMP

OIL

DRAIN

OIL

PUMP

OIL PUMP

DESIGN FLOW RATE = 6.0 TO 6.7 GPM

MAXIMUM PRESSURE = 150 PSI

RATED PRESSURE = 80 PSI

NOTE

OIL JET 1

ROLLER BEARING-SPIRAL

BEVEL PINION SUPPLIES OIL

FLOW TO OIL JETS 3 AND 4

DUPLEX BEARING

OIL FLOW

(ANNULUS - SLOT)

TRIPLEX BEARING

OIL FLOW

(ANNULUS - SLOT)

OIL FILTER DIFFERENTIAL PRESSURE

BY-PASS VALVE AND BY-PASS

INDICATOR BUTTON (ACTUATION

PRESSURE = 14 ±2 PSID AND VALVE

CRACKING PRESSURE = 17 PSID MIN.)

OIL JET 2

MAST BEARING

(ANNULUS-SLOT)

PLANETARY OIL FLOW

SUN GEAR OIL FLOW

(3 JETS)

PLANETARY OIL FLOW

(3 JETS)

SUN GEAR SPLINE

MAIN TRANSMISSION

AND MAST ASSEMBLY

OIL FILTER

OIL FILTER CAP

OIL JET 3 SPIRAL

BEVEL INTO MESH

(NOT SHOWN)

OIL PRESSURE SWITCH

(ACTUATES AT 32 +2 PSI - LOW PRESSURE

GOING DOWN PRESSURE ACTUATES THE LIGHT

AT 38 PSI - HI PRESSURE DEACTUATES THE

LIGHT PRESSURE INCREASING)

OIL PRESSURE TRANSMITTER

TEST PORT

MANIFOLD

TEE FITTING

RESTRICTOR FITTING ASSEMBLY

0.027 TO 0.032 IN. RESTRICTOR HOLE

RESTRICTOR FITTING ASSY

0.035 TO 0.040 IN. RESTRICTOR HOLE

FREEWHEEL INSTALLATION

OIL JET 4 SPIRAL

BEVEL INTO MESH

(NOT SHOWN)


31 JAN 2007—Rev. 2———1-83

1-27-D. TRANSMISSION OILTEMPERATURE AND PRESSUREGAUGE

The transmiss ion oi l temperature andpressure gauge is a dual instrument thatsimultaneously displays oil temperature indegrees Celsius on the right side display andoil pressure in PSI on the left side display.Each side of the indicator is powered by itsown circuit breaker.

The transmission temperature input signal isprovided by a thermobulb installed on thet ransmiss ion o i l f i l te r man i fo ld . Thetransmission oil pressure is provided by atransducer mounted on the transmissiondeck oil manifold.

1-27-D-1. TRANSMISSION OILTEMPERATURE GAUGE

The XMSN OIL TEMP light will be illuminatedwhen the transmission oil temperature switchdetects a temperature of 110 ±5.6°C (230±10°F). The oil temperature switch is mountedon the transmission oil f i l ter manifoldhousing.

1-27-D-2. TRANSMISSION OIL PRESSURELIGHT

The XMSN OIL PRESS light will be illuminatedwhen the transmission oil pressure switchdetects a decreasing oil pressure of 30 ±2 PSI.

Similarly, as the transmission oil pressurebuilds, the oil pressure switch will extinguishthe XMSN OIL PRESS annunciator prior to thepressure passing through 38 PSI.

1-27-E. TRANSMISSION CHIPANNUNCIATOR

The XMSN CHIP l ight wi l l i l luminate i fmagnet ic par t ic les in the o i l systemaccumulate on any one of the three quickdisconnect magnetic chip detectors.

Two magnetic chip detectors, one on theupper case and one on the lower case, aremounted on the main transmission. The lowercase chip detector incorporates a self-sealingvalve which prevents the loss of oil from thegearbox when the chip detector is removed.The third chip detector is mounted on theforward f reewheel hous ing and a lsoincorporates a self-sealing valve whichprevents the loss of oil from the freewheelunit when the chip detector is removed.

1-28. ROTOR SYSTEM

1-28-A. MAIN ROTOR HUB AND BLADES

The rotor assembly (Figure 1-20) is a fourbladed soft-in-plane design with a 35 footdiameter rotor.

The main rotor hub contains a glass/epoxycomposite yoke that acts as a flappingflexure. Elastomeric bearings and damperswhich require no lubrication are utilized. Thehub also incorporates the use of lead-lag,coning/flapping and droop stops.

The main rotor blades are a composite designutilizing a glass/epoxy spar, glass/epoxyskins, and a nomex core afterbody. Theblades incorporate a nickel plated stainlesssteel leading edge erosion strip and arecoated with conductive paint for lightningprotection. The blades are also individuallyinterchangeable.

1-28-B. TAIL ROTOR HUB AND BLADES

The tail rotor (Figure 1-21) is a two-bladedteetering rotor with a 5.42 foot diameter. It ismounted on the left side of the tailboom androtates clockwise when looking inboard fromthe left side of the helicopter.

Teflon l ined pitch change bearings areinstalled in a steel yoke assembly whichutilizes an elastomeric flapping bearing andflapping stops.


1-84———Rev. 2—31 JAN 2007

Figure 1-20. Main Rotor Assembly407_MD_01_0020

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

11.

12.

13.

14.

15.

16.

17.

Cover

Frahm damper

Safety lock

Main rotor blade grip

Yoke

Damper

Elastomeric lead-lag bearing

Pitch link

Mast

Lower cone

Center cone set (if applicable)

Pitch horn

Elastomeric shear bearing

Coning/flapping stop

Upper plate

Mast nut

Upper cone

1

2

3

4

56

14

15

17

16

13

8

7

12

9

11

10

1 Center cone set applicable to S/N 53000 through 53631 Pre TB 407-05-66. S/N 53632 and subsequent

have an integral mast center cone.

NOTE

1


31 JAN 2007—Rev. 2———1-85

Figure 1-21. Tail Rotor Assembly

2

407_MD_01_0021

1211

9

10

1

1

7

SEEVIEW A

VIEW ASTATIC STOP

CROSS SECTION

4

36

5

8

1 S/N 53351 and subsequent, and helicopters modified per ASB 407-99-27 or TB 407-99-17.

DEFORMED

YIELD INDICATOR

0.055 IN. (1.39 mm)

MINIMUM YIELD

INDICATOR DIMENSION

NOTE

1

1.

2.

3.

4.

5.

6.

Tail rotor blade

Tail rotor pitch change link

Crosshead

Counterweight assembly

Tail rotor hub

Pitch horn

7.

8.

9.

10.

11.

12.

Static stop

Tail rotor gearbox

Chip detector

Pitch change mechanism

Breather line

Tail rotor gearbox filler cap


1-86———Rev. 2—31 JAN 2007

The blades are a composite design utilizing aglass/epoxy spar, glass/epoxy skins, and anomex core. The blades incorporate nickelplated stainless steel leading edge abrasionstrip and are coated with conductive paint forlightning protection.

The tail rotor yoke static stop has beendesigned with yield indicators. The yieldindicators provide the ability to visuallydetermine if the tail rotor yoke has beenstressed beyond designed limits. This will beevident by deformation of either of the staticstop yield indicators due to excessive contactwith the yoke (Figure 1-21). If deformation ofeither yield indicator is evident, contactmaintenance personnel prior to further flight.

1-28-C. TAIL ROTOR GEARBOX

The tail rotor gearbox (Figure 1-21), locatedon the aft end of the tailboom, drives the tailrotor. It contains two spiral bevel gearspositioned at 90° angles to the other. The tailrotor gear box has a gear reduction of 2.53 to1.0 which reduces the driveshaft input speedof 6317 RPM to an output shaft speed of 2500RPM.

The gearbox has a se l f conta ined o i llubrication system, non vented filler cap, anda magnetic chip detector.

1-28-D. TAIL ROTOR CHIP LIGHT

The T/R CHIP light will illuminate if magneticpart ic les in the oi l accumulate on themagnetic chip detector.

The chip detector incorporates a self-sealingvalve which prevents the loss of oil from thegearbox when the chip detector is removed.

1-29. ROTOR SYSTEM INDICATORS

1-29-A. DUAL TACH GAUGE

The Dual Tachometer has two pointers thatsimultaneously display Rotor RPM (NR) on theouter scale and the engine NP on the innerscale. All scales are in percent RPM.

The NP side of the instrument is powered bythe 28 VDC bus through its own circuitbreaker.

An NP Speed Pickup mounted on the engineprovides two separate signals to the FADEC/ECU. One signal is the primary driver of theNP indicator and is also a secondary signalfor the FADEC/ECU. The second signal is theprimary input to the FADEC/ECU. The primarysignal to the indicator will be passed throughthe FADEC/ECU even when the FADEC/ECU isnot powered.

The NR side of the instrument is powered bythe 28 VDC bus through its own circuitbreaker.

An N R speed p ickup mounted on thetransmission lower case provides threeseparate identical signal outputs of rotorRPM. One signal is sent to the NR circuit ofthe dual tachometer gauge. One signal is sentto the FADEC/ECU and the other is sent to theRotor RPM sensor switch.

1-29-B. RPM LIGHT AND WARNINGHORN

The RPM light and warning horn circuit isdesigned to activate when the main rotor RPM(NR) is less than 95%. The RPM light will alsobe illuminated at a NR speed of 107% andhigher.

1-29-B-1. NR LESS THAN 95%

If the rotor RPM sensor detects a main rotorRPM (NR) of less than 95%, it will illuminatethe RPM light and (continuous sounding) lowrotor RPM horn. The low rotor RPM horn canbe muted by pressing the warning horn muteswitch.

1-29-B-2. NR 107% OR GREATER

If the rotor RPM sensor detects a main rotorRPM NR of 107% or greater, it will illuminatethe RPM light. The low rotor RPM horn will notbe activated.

31 JAN 2007 Rev. 2 1-87


1-30. FLIGHT CONTROL SYSTEM

1-30-A. ROTOR CONTROLS

Main rotor and tail rotor flight control systems(Figure 1-22), consisting of cyclic, collectiveand anti-torque controls are used to regulatethe helicopter attitude, altitude and directionof flight. The flight controls are hydraulicallyboosted to reduce p i lo t e f for t and tocounteract control feedback forces.

1-30-B. MAIN ROTOR

Main rotor cyclic and collective flight controlsregulate pitch and roll attitude and thrust.Control inputs from the cyclic and collectivecontrol sticks (Figure 1-22) in the cockpit aretransmitted by push-pull tubes to hydraulicservo actuators mounted on the top deck. Theactuators operate the cyclic and collectivelevers, which ra ise, lower, and t i l t theswashplate. The swashplate converts fixedcontrol inputs to the rotating controls andallows cyclic and collective pitch inputs to themain rotor.

In the case of loss of hydraulic pressure to theservo actuators, springs are installed inparallel to the cyclic and collective push-pulltubes on the cabin roof to assist the pilot withthe increased control feedback forces.

1-30-B-1. CYCLIC

The cyclic control stick (Figure 1-22) ismounted under the pilots crew seat andprotrudes from the forward bulkhead of thecrew seat. The fore and aft cyclic input isconnected through push-pull tubes to thecyclic hydraulic servo actuators. In addition,all cyclic fore and aft movement is fed througha cam assembly that automatically adds anamount of lateral cycl ic input that is apercen tage of the fo re and a ft cyc l icmovement. A spring canister is provided inline with the cam input to permit cyclic

movement in the event that the cam assemblybecomes jammed.

The lateral cyclic input is connected throughpush pull tubes to the cyclic hydraulic servoactuator. The hydraulic servo actuatorsoperate bellcranks and push pull tubes that tiltthe swashpla te non-rotat ing r ing . Theswashplate rotating ring tilts likewise andactuates the pitch links which control theplane of rotation of the main rotor.

The cyclic control stick grip contains a two-position intercommunication/radio transmitswitch and a cargo hook release switch. Anadjustable friction control knob, located at thebase of the cyclic stick where it protrudesthrough the forward crew seat bulkhead,allows the pilot to set the desired amount ofcontrol stiffness for flight or to lock the cycliccontrol stick during ground operation orshutdown.

A cyclic stick position switch (cyclic centeringswitch) is at tached to the cycl ic st ickbellcrank. When the helicopter is on theground, this switch will cause the CYCLICCENTERING annunciator in the caution/warning panel to illuminate when the stick isnot centered.

1-30-B-2. COLLECTIVE

The collective control stick (Figure 1-22) ismounted between the pilot and copilot crewseats. The collective control stick controls thecollective hydraulic servo actuator throughpush-pull tubes. This operates the collectivelever mounted on the top of the transmission.The collective lever raises and lowers theswashplate ball-sleeve assembly and thecyclic levers to induce collective pitch to themain rotor blades without affecting the cyclicpath. A spring is installed under the copilot'screw seat to balance the required force toraise and lower the col lect ive with thehydraulic boost system operating.

1-88 Rev. 2 31 JAN 2007


Figure 1-22. Flight Controls

MAIN ROTOR

PITCH LINK

SWASHPLATE

ASSEMBLY

NON ROTATING RING

ROTATING RING

SEEDETAIL B

SEEDETAIL A

ALTERNATE

PEDALS

407_MD_01_0022

1

19

2018

34

7

5

2

13

8

14

17

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

11.

12.

13.

14.

15.

16.

17.

18.

19.

20.

Pilot tail rotor control pedal assembly

Pilot cyclic stick

Cyclic friction knob

Cyclic lateral balance spring

Cyclic centering switch

Cam override spring

Pilot collective stick

Collective friction knob

Collective servo actuator

Cyclic servo actuators

Balance springs

Tail rotor servo actuator

Collective balance spring

Copilot collective stick

Collective pitch transducer (FADEC SYS)

Cyclic cam

Cyclic longitudinal balance spring

Copilot cyclic stick

Copilot tail rotor control pedal assembly

Control pedal spring

16

12

10

9

11

15

6

LDGLTS

BOTH

FWD

OFF

DISENG

START

FLOATINFLATE

FLOATARM

DETAIL APILOT COLLECTIVE STICK

DETAIL BCYCLIC GRIP

TYPICAL

11

1 Alternate pedals installed in production at S/N 53730 and subsequent.

NOTE

1


31 JAN 2007—Rev. 2———1-89

A collective friction knob is located near thebase of the collective stick between the pilotand copilot seats. A throttle twist grip for theengine is mounted on the collective stick. Amechanical idle release push button islocated in front of the twist grip throttle. Aswitchbox located on the forward end of thecollective stick provides a base for the enginestart switch and landing light switch.

1-30-C. TAIL ROTOR

The tail rotor, or anti-torque, flight controls(Figure 1-22) provide pitch adjustment of thetail rotor blades for yaw control. A set ofpedals on the cockpit floor, forward of thepilot seat, are connected to a directionalcontrol hydraulic servo actuator, located inthe aft fuselage near the tailboom. Push-pulltubes connect the actuator to the fixed pitchchange mechanism of the tail rotor gearbox.The tail rotor fixed mechanism is connectedto the rotating controls through a rotatingpush-pull tube. The push-pull tube attaches toa sliding crosshead that moves in and out onsplines on the tail rotor mast to provide pitchcontrol. Rotating counterweights minimize thecontrol forces required.

The tai l rotor control pedals contain abellcrank pedal adjuster, which provides formanual ad justment o f pedal pos i t ionaccording to the pilot’s needs. Alternatepedals may also be installed which allow thepilot to manually adjust the position of thepedal foot rests.

For helicopters with dual controls, the copilotfully functional tail rotor control pedalassembly is installed on the floor in front ofthe copilot seat to provide a means for thecopilot to control the tail rotor assembly. Thecontrol pedals are linked to the pilot pedals bymeans of control tubes and a bellcrank. Thecopilot pedals can also be positioned, asdesired, by means of the pedal adjuster.Alternate pedals may also be installed whichallow the copilot to manually adjust theposition of the pedal foot rests.

1-30-D. AIRSPEED ACTUATED PEDALSTOP SYSTEM

The Airspeed Actuated Pedal Stop system iscomprised of a Pedal Restrictor Control Unit(PRCU), an actuat ing rotary solenoid,positioning sensing microswitch, and apress-to-test PEDAL STOP (PTT) switch.

A PEDAL STOP segment (Figure 1-4) on theCaut ion /Warn ing pane l ind icates anymalfunction of the system. System power isprovided through a 5-amp circuit breakerlocated in overhead console (Figure 1-5).

The helicopter pitot and static installationinter faces w i th the PRCU. The PRCUcalculates airspeed from the pitot and staticinputs and when greater than 55 ±5 KIAS,drives the solenoid to extend the pedal stoprestrictor into the left pedals range of travel.The PRCU will trigger the solenoid to retractthe pedal stop when calculated airspeed fallsbelow 50 ±5 KIAS.

Upon full extension of the pedal stop, theposition sensing microswitch is activated andthe PRCU illuminates the ENGAGED messageon the PEDAL STOP (PTT) switch. Thismessage is extinguished when the pedal stopis retracted.

The Airspeed Actuated Pedal Stop Systemdiagram is shown in Figure 1-23.

1-31. HYDRAULIC SYSTEM

The hydraulic system (Figure 1-24) providesboost power for the cyclic, collective andanti-torque fl ight controls. The systemincludes a pump, reservoir, pressure andreturn filter assemblies, pressure and returnmanifold, pressure monitoring sensor,solenoid valve, pressure relief valve, flightcontrol servo actuators, and interconnectingtubing and fittings.

The hydraulic pump is mounted on and drivenby the transmission. The pump is a variablede l ivery pressure compensa ted ,


1-90———Rev. 2—31 JAN 2007

self-lubricated type designed to operatecontinuously and provide a rated dischargepressure of 1000 -25/+50 PSI. Hydraulic fluidis supplied to the pump from a vented gravityfeed reservoir mounted forward of thetransmission. Fluid passes through the pump,a pressure filter, and a solenoid valve. Thesolenoid valve is controlled by a HYD SYSswitch located on the overhead console.When HYD SYS switch is ON, the solenoidvalve is de-energized to the open position andfluid is routed to cyclic, collective, andanti-torque actuators. From the actuators,fluid passes through a return filter to thereservoir. Both the pressure and return filterassemblies have filter indicators. Theseindicators are activated when the differentialpressure across the filter is 70 ±10 PSI. Thepressure filter does not contain a bypassvalve. The pressure filter will clog completelyin order to prevent contaminated hydraulicfluid from being pumped through the system.The return filter contains a bypass valve thatwill allow fluid to bypass the filter if it sensesa differential pressure of 100 ±25 PSI acrossthe filter. When the differential pressuredecreases to 60 PSI, the bypass valve willclose. This allows fluid returning from theservo actuators to return to the reservoir evenif the return filter is completely clogged.

A relief valve is incorporated in the system,between the pressure filter and the solenoidvalve. The relief valve is normally closed.When the pressure reaches 1225 ±150 PSI, therelief valve will open to protect the systemfrom damage. The relief valve will reset whenpressure drops to approximately 1075 PSI.

1-31-A. HYDRAULIC INDICATORS

Hydraul ic system indicators include aHYDRAULIC SYSTEM caution light on thecaution/warning panel and filter bypassindicators located on both the pressure andreturn filter assemblies.

1-31-B. HYDRAULIC FILTERINDICATORS

Each filter assembly contains a filter indicatorthat indicates an impending clogged filter.The indicator consists of a red buttonmounted on the filter assembly housing.When the differential pressure across thefilter is 70 ±10 PSI, the red button will rise. Toprevent inaccurate indications of bypass, theindicator will not work when the hydraulicfluid temperature is less than 35°F (2°C). Ifhydraulic fluid temperature is more than 35°F(2°C) , the ind icator g ives the correctindication of clogging, even if the ambienttemperature is below 35°F.

If a filter button is extended, indicating animpending clogged filter, maintenance actionshould be performed prior to next flight. Thefilter indicator is reset by pushing the button in.

1-31-C. HYDRAULIC SYSTEM LIGHT

The HYDRAULIC SYSTEM l ight circuit isdesigned to illuminate when the hydraulicsystem fluid pressure is below the minimumlimit.

The hydraulic system light will be illuminatedwhen the hydraulic pressure switch detects adecreasing pressure of 650 -0/+100 PSI, andwill extinguish on an increasing pressure at750 +0/-100 PSI. The hydraulic pressureswitch is mounted on the hydraulic manifold,forward of the hydraulic actuator support.

1-32. ELECTRICAL SYSTEM

The helicopter is equipped with a 28 VDCelectrical system (Figure 1-25). Power for thissystem is obtained from a nickel-cadmium 24volt, 17 amp/hour battery or optional 24 volt,28 amp/hour battery and a 30 volt, 200-ampstarter-generator. The starter-generator hasbeen derated to 180 amps to ensure adequatecooling under all operating conditions up to18,000 feet Hp. Refer to the BHT-407-FMlimitations for operations above 18,000 feetHp.

31 JAN 2007 Rev. 2 1-91


Figure 1-23. Airspeed Actuated Pedal Stop System407_MD_01_0023

28 VDC PWR

TEST INPUT

POSITION ANN GND

FAULT ANN

SOLENOID DRIVER

POSITION IND

POSITION IND RTN

POWER GND

FAULT ANN GND

LTG POWER28/15 VDC

LTG POWER28/15 VDC

PEDAL STOP (PTT)SWITCH

PEDAL STOPPTT

CAUTIONWARNING PANEL

PEDAL STOPPTT

ENGAGED

6540CB1

PRCU

PITOTINPUT

STATICINPUT

1-92 Rev. 2 31 JAN 2007


Figure 1-24. Hydraulic System407_MD_01_0001

17.

18.

Tail rotor servo actuator

Hydraulic pressure switch

RETURN

SEQUENCEVALVE CHECK

VALVES

INPUT FROM

FLIGHT CONTROLS

CYLINDER

OUTPUT

SERVO ACTUATOR - TYPICALSEE DETAIL B

PRESSURERETURNSUCTION

OPERATING PRESSURE 1000 PSI -25/+50

HYDRAULIC FLUID MIL-H-5606

PRESSURERETURN

TEST PORT

DIFFERENTIAL

RELIEF VALVE

HYDRAULIC SYSTEM

CAUTION LIGHT

CYCLIC18

15 15

16 171

23

4

5 6

7

8

9

10 11

12

13

14

85

CYCLIC COLLECTIVE TAIL ROTOR

SEE DETAIL B

SEE DETAIL B

SEE DETAIL B

SEE DETAIL A

SERVO

ACTUATOR

PRESSURE

RETURN

ENERGIZED-

SYSTEM OFF

DE-ENERGIZED-

SYSTEM ON

SOLENOID VALVE SCHEMATIC

DETAIL A

SERVO

ACTUATOR

PRESSURE


31 JAN 2007—Rev. 2———1-93

Figure 1-25. DC Electrical System407_MD_01_0024

TYPICAL

BATTERY

BATTERY

RELAY

EXTERNAL

POWER RELAY

EXTERNAL

POWER RECEPTACLE

STARTER

RELAY

GENERATOR

RELAY

FADEC/START

RELAY

FADEC

MODE SWITCH

FADEC/ECU

START (-)

AUTO

MANUAL

GEN FIELD

10 15

GEN FIELD

AMMETER

SHUNT

START

SWITCH

START

DISENG

AUTO

MANUALENG START

5

STARTER

GENERATOR

DC CONTROL UNIT/

VOLTAGE REGULATOR

GEN RESET

GENERATOR

SWITCH

RESET

OFF

GEN

C

B

5

A

E

D

L

D

K

H

M

J

G

BATTERY

SWITCH

OFF

BAT (ON)

AVIONICS BUS

BUS BAR BUS BAR

AVIONICS SWITCH

28

VD

C B

US

- +

- + +

+

-


1-94———Rev. 2—31 JAN 2007

Major components of DC power systeminclude battery, starter-generator, DC controlunit, voltage regulator, relays, 28 VDC bus,and circuit breakers. All circuits in electricalsystem are single wire with fuselage commonground re tu rn . Negat ive termina ls o fstarter-generator and battery are grounded tohelicopter structure. Controls for electricalsystem are located on overhead console andinstrument panel.

Generator is provided with over voltage,under voltage and reverse current protection.If an over voltage (32 ±0.5 VDC), under voltage(18 ±1.8 VDC), or reverse current (0.08 to 0.150VDC for 350 milliseconds) is detected, the DCcontrol unit/voltage regulator will disconnectthe generator from the system. Failure of thegenerator can be determined by GEN FAILlight, a zero ammeter reading, and batteryvoltage displayed on the voltmeter.

In the event power from the generator is lost,emergency power available from the batterycan be maximized by pulling the circuitbreakers on all non-essential systems. In theevent of a total electrical failure (hard short onbus), if the battery switch is immediatelyplaced to the OFF position, the 17 amp/hourbattery, assuming 80% charged, can supplyfuel boost and transfer pumps for a period ofapproximately 1.7 hours. The optional 28 amp/hour battery, assuming 80% charged, cansupply fuel boost and transfer pumps for aperiod of approximately 2.8 hours.

1-32-A. EXTERNAL POWER

External power may be supplied to thehelicopter by means of a receptacle locatedon the lower front section of helicopter. 28VDC Ground Power Unit (GPU) shall be 500amps or less to reduce risk of starter damagefrom overheating.

If external power was used to power the startand the battery switch was left in the OFFposition, it is important to position the batteryswitch to ON prior to removing the externalpower source (refer to BHT-407-FM-1, Normal

Procedures). If all sources of electrical powerare removed from the ECU with the engine atidle in AUTO mode, the start solenoid valve inthe HMU will open, causing the engine todecelerate and possibly flame out. If theBattery switch is inadvertently left OFF andthe external power source is removed, do notattempt to reapply power when a decrease inNG speed is noted. Thrott le should beposit ioned to cutoff . Reappl icat ion ofe lectr ica l power could cause an overtemperature condition due to the reduced NGspeed and reintroduction of fuel by theFADEC system.

1-32-B. BATTERY SWITCH

The battery switch is insta l led on theoverhead console and controls the batteryrelay which connects the battery to the DCbus. The switch has two positions OFF andBATT.

In the event BATTERY HOT light comes ON,BATT switch shall be turned OFF. If BATTswitch is turned OFF and BATTERY RLY lightilluminates, this indicates battery relaycontacts have not opened. If battery relaycontacts have not opened, battery wil lcont inue to receive a charge from thegenerator. To prevent this, pilot should turnGEN switch to OFF. Battery will continue torun all electrical systems through the closedbattery relay. To reduce load on overheatedbattery, pilot should pull circuit breakers onall non-essential systems. Left FUEL XFR/BOOST circuit breaker switch should be leftON to insure fuel transfer from the forwardfuel tank to the main fuel tank continues.

The BATT switch also configures the DCpower feed to the left fuel boost pump and theleft fuel transfer pump between the DC bus(battery switch positioned to BATT) and thehelicopter battery (BATT switch positioned toOFF).

In the event the battery switch is positioned toOFF dur ing he l icopter operat ions , analternate circuit (Figure 1-17) is provided to


31 JAN 2007—Rev. 2———1-95

allow operation of the left fuel transfer and leftfuel boost pumps. During this condition, withthe FUEL VALVE switch positioned to ON,battery voltage is supplied through the fuelboost/xfr backup circuit breaker, the fuelvalve switch, the battery switch, and the leftfuel xfr/ boost circuit breaker switch to the leftfuel transfer and left fuel boost pumps.

1-32-C. BATTERY CHARGING (17 AND 28AMP/HOUR BATTERIES)

As a maintenance function, battery chargingmay be accomplished with battery installed inhelicopter, due to minor depletion, using aGPU. The GPU must incorporate a goodquality constant voltage regulator, a variablevoltage selector and an amperage indicator.

NOTE

It is recommended that the chargingprocedure not exceed 45 minutes indurat ion. Frequent use of th isprocedure may cause loss ofelectrolyte due to gassing throughelectrolysis. Reduced levels ofe lec t ro ly te can cause ce l limbalances and lower overall batterycapacity. This procedure is notin tended to take the p lace o fscheduled battery maintenanceprocedures.

CAUTION

BATTERY HOT CAUTION LIGHTMUST BE CONTINUOUSLYMONITORED DURING THISPROCEDURE. IF BATTERY HOTLIGHT ILLUMINATES, BATTERYSWITCH SHALL BE SET TO OFF ANDGPU POWER REMOVED TO REDUCEPOSSIB IL ITY OF THERMALRUNAWAY.

Battery switch may be set to ON after GPUpower is applied and voltage adjusted to 28.5VDC (do not exceed 28.5 volts). Battery

charging is supplied by the GPU and must bemonitored. Charging will be completed whenGPU output indicates approximately 8 amps.In the event BATTERY HOT warning lightilluminates, battery shall be turned OFF andGPU power removed.

1-32-D. GENERATOR SWITCH

The generator switch is installed on theoverhead console and controls generatoroutput by opening and closing the generatorfield circuit. The switch is a double pole,double throw, spring loaded design with onlymomentary contact in the RESET position.The switch has 3 positions, GEN, OFF, andRESET.

With the generator switch positioned to GEN,its function is to complete the generator fieldcircuit between the starter generator and thegenerator control unit/voltage regulator.Under normal operating conditions, this willallow the generator control unit/voltageregulator to monitor and control the outputvoltage of the starter generator and in turnconnect the output of the generator to the 28VDC bus through the generator relay.

Positioning the generator switch to OFFopens the generator field circuit whichremoves control of the generator control unit/voltage regulator from the generator andgenerator relay. The generator relay will open,removing the generator from the 28 DC bus.

In the event an over voltage condition isdetected by the generator control unit/voltageregulator, an internal regulator trip relaycircuit will be activated. Positioning thegenerator switch to RESET provides buspower, from the battery, to the generatorcontrol unit/voltage regulator which will resetthe in terna l t r ip re lay c i rcu i t . I f themalfunction condition persists following theRESET, further attempts to reset should notbe made.

Additionally, following a start using a GPU,ensure battery switch is positioned to ON and


1-96———Rev. 2—31 JAN 2007

GPU is disconnected prior to positioning thegenerator switch to GEN. Positioning thegenerator switch to GEN with the GPUconnected may cause a reverse currentsituation and trip the generator off line.

1-32-E. START SWITCH

The start switch is located on the collectiveswitch box. It contains two sets of springloaded contacts which provide momentarycontact in either the START or DISENGpositions. When the switch is positioned tostart or disengage, only one set of contactsmove (Figure 1-25). For a description of aSTART IN AUTO MODE or START IN MANUALMODE refer to paragraph 1-12-D or paragraph1-12-F.

1-32-F. ELECTRICAL SYSTEMINDICATORS

Electrical system indicators include a DCammeter, voltmeter, BATTERY HOT light,BATTERY RELAY light, START light, and GENFAIL light.

1-32-G. FUEL PRESSURE/DC AMMETER

The fuel pressure/ammeter gauge is a dualdisplay instrument. The ammeter indicatesthe load in amperes that is being supplied tothe 28 VDC bus by the engine dr ivengenerator.

The ammeter side of the indicator is poweredby its own circuit breaker.

1-32-G-1. AMMETER — 200 AMPS MAXSCALE

For ammeter readings above 200 amps(maximum scale) the indicator will indicate asfollows:

Amperes 201 to 300 amps — Maintain fullscale for 2 minutes. After 2 minutes, dropto zero.

Amperes 301 to 400 amps — Maintain fullscale for 5 seconds. After 5 seconds,drop to zero.

If ammeter indication drops to zero,pos i t ion generator swi tch to OFF.Following a brief time, generator switchcan be positioned to GEN (ON). Refer tothe BHT-407-FM-1 for generator loadlimitations.

1-32-G-2. AMMETER — 400 AMPS MAXSCALE

Ammeter indications will be continuousregardless of load. Indicator will not drop tozero if limits are exceeded. Refer to theBHT-407-FM-1 for generator load limitations.

To ensure limits are not exceeded, pilot canswitch generator to OFF prior to generatorexceedance being reached. Following a brieftime, generator switch can be positioned toGEN (ON).

1-32-H. VOLTMETER

The voltmeter is included in a multifunctionindicator mounted in the upper left area of theinstrument panel. The indicator also displaysOutside Air Temperature and Clock functions.

A button located on the center top of theinstrument changes the top display betweenvolts (e.g. , 28E) and OAT (Celsius andFahrenheit). When power is applied to theinstrument the display defaults to thevoltmeter reading.

The voltmeter display receives its power fromthe 28 VDC bus through the OAT/V INSTRcircuit breaker. The 28 VDC bus through thiscircuit breaker is also the source of thevoltage value read and displayed by thevoltmeter. When all electrical power is turnedoff, the voltmeter display disappears.


31 JAN 2007—Rev. 2———1-97

1-32-I. BATTERY HOT ANNUNCIATOR

The BATTERY HOT light is utilized with eitherthe basic ship 17 amp/hour or optional kit 28amp/hour battery installed. The 17 amp/hourbattery incorporates two thermal switcheswhile the 28 amp/hour battery incorporatesthree thermal switches. While both of the 17amp/hour battery thermal switches are usedin the BATTERY HOT light circuit, only two ofthe three thermal switches available on the 28amp/hour battery are used.

The 17 amp/hour battery thermal switches willclose at a temperature of 145 ±5°F (62.7±2.8°C). The 28 amp/hour battery switches willclose at a temperature of 160 ±5°F (71.1±2.8°C).

When any one of the thermal switches closes,the BATTERY HOT light will be ON.

1-32-J. BATTERY RELAYANNUNCIATOR

The BATTERY RELAY light will be ON if thebattery relay has remained in the closed(energized) position after the battery switchhas been set to OFF.

If the battery relay remains energized after thebattery switch has been set to OFF, batterypower will remain on the 28 VDC bus. This willpower the caution and warning panel to allowillumination of the BATTERY RELAY light,even if the generator is off.

1-32-K. START ANNUNCIATOR

The START light will be illuminated when thestarter relay is energized. The starter relay willbe energized when the start switch ispositioned to START as follows:

With the FADEC MODE switch positionedto AUTO, the starter relay will stayengaged until the gas producer (NG)speed reaches 50 ±1%.

With the FADEC MODE switch positionedto MANUAL, the starter relay will stay

engaged until the start switch is releasedfrom the START position.

1-32-L. GEN FAIL ANNUNCIATOR

The GEN FAIL light is controlled by generatorrelay. The light will be ON when the generatorrelay is de-energized and not connecting thegenerator output to the DC bus.

The generator relay will be energized by thegenerator control unit/voltage regulator whengenerator output climbs through a thresholdof 24 ±2.4 VDC. Prior to the generator relaybeing energized, the GEN FAIL light will beON. Once the generator relay is energized, theGEN FAIL light will be OFF.

1-33. PITOT STATIC SYSTEM

The pitot-static system (Figure 1-26) utilizes aconventional impact air and ambient airpressure sensing system.

The pitot-static system consists of a heatedpitot tube and right and left heated static portsand interconnecting tubing.

The pitot tube supplies impact air pressure tothe airspeed indicator and PRCU (paragraph1-30-D). The two static ports are connectedtogether to equalize the static pressure, whichis supplied to the airspeed indicator, altimeter,vertical speed indicator, and PRCU.

1-34. BASIC FLIGHT INSTRUMENTS

The basic set of flight instruments includes anairspeed indicator, pressure altimeter, verticalspeed indicator, and inclinometer.

1-34-A. AIRSPEED INDICATOR

The airspeed indicator presents airspeed from0 to 150 knots. The indicator is scaled in 20knot increments from 0 to 20 knots and in 5knot increments from 20 to 150 knots. Amaximum speed red line is located at 140knots and a maximum autorotation speed red/white line is located at 100 knots.


1-98———Rev. 2—31 JAN 2007

Figure 1-26. Pitot Static System407_MD_01_0025

1.

2.

3.

4.

5.

6.

7.

Altimeter

Vertical speed indicator

Airspeed indicator

Pedal restrictor control unit (PRCU)

Pitot static drains

Static ports (heated)

Pitot tube (heated)

1

2

3

5

6

6

7

4


31 JAN 2007—Rev. 2———1-99

1-34-B. ALTIMETER

The pressure altimeter presents an altitudereading in feet above mean sea level (MSL)based on the relationship between the staticair pressure and the barometric setting on thealtimeter. The barometric setting may beadjusted to reflect the current barometricpressure corrected to sea level in inches ofmercury or millibars.

1-34-C. VERTICAL SPEED INDICATOR(VSI)

The vertical speed indicator presents rate ofclimb or descent from 0 to 4000 feet perminute. A maximum rate of climb red line islocated at 2000 FPM.

1-34-D. INCLINOMETER

The inclinometer consists of a curved glasstube, bal l , and damping f luid. The ballindicates the directional balance of thehelicopter. If the helicopter is in a slip or askid, the ball will move off center.

1-34-E. OPTIONAL FLIGHTINSTRUMENTS

Optional flight instruments include an attitudeindicator, directional gyro, turn-and-slipindicator and encoding altimeter.

1-34-F. ATTITUDE INDICATOR

The attitude indicator is powered electricallyand presents pitch and roll attitudes of thehelicopter in relation to the horizon. A fail flagwill be visible when power to the indicator islost warning attitude information is unreliable.

The manual caging knob should be pulled outto erect the gyro prior to positioning the ATTswitch to ON. The pitch trim knob is used toadjust the helicopter symbol vertically.

1-34-G. DIRECTIONAL GYRO

The DG indicator is electrically powered. TheDG is not slaved to a magnetic flux valve andtherefore, must be set prior to takeoff andperiodically during fl ight to correct formagnet ic va r ia t ion and gyroscopicprecession. A fail flag will be visible whenpower to the indicator is lost warningdirectional information is unreliable.

1-34-H. TURN-AND-SLIP INDICATOR

The turn-and-slip indicator is poweredelectrically and has a pointer that indicatesdirection and rate of turn and an inclinometerthat indicates a coordinated or skidding orslipping turn. A red fail flag will be visiblewhen power to the indicator is lost, warningturn information is unreliable.

1-34-I. ENCODING ALTIMETER

The ENCODING altimeter presents an altitudereading in feet above mean sea level (MSL)based on the relationship between the staticair pressure and the barometric setting on thealtimeter. The barometric setting may beadjusted to reflect the current barometricpressure corrected to sea level in inches ofmercury or millibars.

Additionally, the encoding altimeter providescoded pulses to the transponder and GPS (ifinstalled) for altitude reporting.

1-35. NAVIGATION SYSTEMS ANDINSTRUMENTS

The basic navigation equipment consists of amagnetic compass.

1-35-A. MAGNETIC COMPASS

The magnetic compass is a standard, nonstabilized, magnetic type instrument mountedon a support which is attached to the rightside of the forward crew cabin. The compassis used in conjunction with the compasscorrection card.


1-100———Rev. 2—31 JAN 2007

1-35-B. OPTIONAL NAVIGATIONEQUIPMENT

Optional navigation equipment currentlyavailable through Bell Helicopter consists of ahorizontal situation indicator (HSI), globalpositioning system (GPS), VOR indicator,automat ic d i rect ion f inder (ADF) , andtransponder. Basic operational information isprovided for each installation. For expandedinformat ion , re fer to the equipmentmanufacturer’s operating manual.

1-35-B-1. HORIZONTAL SITUATIONINDICATOR (HSI)

The optional Kl-525A horizontal situationindicator (HSI) (Figure 1-4) provides a displayof the gyro stabilized magnetic heading andhelicopter position relative to the VOR andlocalizer and glideslope beams. Globalpositioning system (GPS) course information,if applicable, may also be displayed on theHSI. Switching between NAV or GPS isselected through a NAV/GPS switch.

Adjustments on the HSI include a headingselect knob for setting the desired magneticheading, and a course select knob for settingthe desired course. The course deviation barindicates displacement from the selectedradial, track, or ILS localizer beam. Theglideslope pointer indicates displacementfrom the ILS glideslope beam. An ambiguitypointer indicates whether the selected coursewill lead TO or FROM the station. The NAV failflag warns the pilot of weak or unreliable VOR/ILS/GPS navigation signals and the HDG flagwarns of gyrocompass failure.

The Bendix/King KCS 55A Gyromagneticcompass system is powered from the 28 VDCbus and prov ides magnet ic head inginformation, which is displayed on the HSI.The system includes a KG-102A slaveddirectional gyro, KMT-112A magnetic azimuthtransmitter, KA-51B slaving accessory, and aKl-525A HSI indicator.

The KA-51B slaving accessory control panel,located on the pedestal, provides two toggleswitches and a slaving indicator. The SLAVE/FREE switch engages the slave gyro modewhen in SLAVE position and places the gyroin the free mode when in the FREE position.The CW/CCW switch is a momentary switchwhich p rov ides c lockwise andcounterclockwise manual slewing when thesystem is in the free gyro mode.

A slaving meter on the control panel indicatesthe instantaneous error between the compasscard presentation and the signal from the fluxvalve.

1-35-B-2. GLOBAL POSITIONING SYSTEM(GPS)

GPS is a Department Of Defense (DOD)operated global coverage, satellite-basednavigation system.

The optional KLN 89B GPS (Figure 1-4) is along-range, GPS based RNAV System. Itconsists of a panel-mounted receiver/displayunit and one antenna. The system is designedto supply the Pilot with navigation guidanceand position information in three-dimensions:latitude, longitude, and altitude.

The NAV/GPS push button on the instrumentpanel (if installed) will configure the HSI todep ic t VOR/LOCALIZER/GLIDESLOPEinformation when selected to NAV and GPSinformation when selected to GPS.

Additionally, pressing the caution panel testbutton will turn on the NAV/GPS switch lights.

Refer to appropriate FMS and the Bendix/King(Allied Signal) KLN 89B Pilots Guide foradditional information.

1-35-B-3. VOR INDICATOR

An optional Kl-208 course deviation indicator(Figure 1-4) may be installed as a primaryVOR/LOCALIZER indicator.


31 JAN 2007—Rev. 2———1-101

An omni bearing selector (OBS) and TO/FROM ambiguity pointer are provided. TheNAV fail flag warns of weak or unreliable VOR,localizer signals, or equipment malfunction.

1-35-B-4. AUTOMATIC DIRECTION FINDER(ADF)

The optional automatic direction finder set(Figure 1-4) includes a KR 87 ADF receiver, Kl227 ADF indicator, and a KA 44B loop/senseantenna. The receiver includes active/standbyfrequency selection and a flight or elapsedtimer.

The right-hand side of the KR 87 receiver willdisplay either the standby frequency or theflight or elapsed timer. If it is in one of thetimer modes, press the FRQ button once todisplay the standby frequency. Press the FRQbutton again to exchange the active andstandby frequencies.

Press the FLT/ET button to return to one ofthe timer modes. The unit will enter whichevert imer mode was displayed previous topushing the FRQ button. Therefore, if it entersthe ET (elapsed timer) mode, press the FLT/ETbutton once more to enter the FLT (flighttimer) mode. The flight timer should reset tozero if the unit is turned off and turned backon.

The elapsed timer (ET) has two modes:count-up and count-down. If the display is inthe FLT mode, press the FLT/ET button onceto display the elapsed timer. When power isapplied, the ET is in the count-up modestarting at 0. When in the count-up mode, thetimer may be reset to 0 by pressing theRESET button.

To enter the count-down mode, hold theRESET button in for approximately 2 secondsuntil the ET message begins to flash. Thedisplay may now be set to any time up to 59minutes and 59 seconds. The timer willremain in this ET set mode (whenever ETmessage is flashing) for 15 seconds after anumber is preset, or until the RESET, FLT/ET,

or FRQ button is pressed. The preset numberwill remain unchanged, until the RESETbutton is pressed, at which time it will beginto count down. When the time reaches 0, itwill begin counting up from 0 and the displaywill flash for 15 seconds (regardless of thecurrent display). While the elapsed timer iscounting down, pressing the RESET buttonwill have no effect unless it is held for 2seconds, putting the timer into the ET setmode.

Pressing the FLT/ET button will exchange thetwo timers in the display, or will cause the lasttimer that was displayed to reappear if thestandby frequency is being displayed.Pressing the FRQ button will cause thestandby f requency to reappear, andsubsequent actuation will cause the activeand standby frequencies to be exchanged.

1-35-B-5. ATC TRANSPONDER

The optional KT-70 or KT-76A Transponder(Figure 1-4) is a radio transmitter and receiveroperating at 1090 MHz in transmit mode and1030 MHz in receive mode. The equipment isdesigned to fulfill the role of airborne beaconunder the requirements of the Air TrafficControl Radar Beacon System (ATCRBS).Range and azimuth are determined by thetransponder pulsed return in response tointerrogation from the ground radar site. Anidentity code number, selected at the frontpanel, is transmitted at a Mode A reply. Formode C altitude reporting capability, thetransponder must be used in conjunction witha reporting (encoding) altimeter and operatedin ALT mode. After pressing the IDENT buttonwhen interrogated, the transponder willt ransmi t a spec ia l pu lse caus ing theassociated PIP to bloom on the ATC display.Either transponder can reply on any of 4096preselected codes.

1-35-B-6. VHF NAV/COMM SYSTEM

The op t iona l VHF NAV/COMM system(Figure 1-4) can include a KX-155 NAV/COMM,


1-102———Rev. 2—31 JAN 2007

or KX-165 NAV/COMM as COMM 1 andKY196A COMM 2.

The KX-155 and KX-165 are VHF NAV/COMMtransceivers which operate within thefrequency range of 118.00 MHz to 136.975MHz, in 25 kHz increments (760 channels) forCOMM, and 108.00 MHz to 117.95 MHz, in 50kHz increments (200 channels), for NAV.

Both the KX-155 and KX-165 NAV/COMM havetwo displays: the left-hand display for COMMfrequencies and the right-hand display forNAV frequencies. The COMM display presentsa USE and STANDBY frequency with a Tappearing between the frequency numbers toindicate operation in the transmit mode. TheNAV display presents a USE and STANDBYfrequency. For both COMM and NAV displays,the desired frequency is entered into theSTANDBY window and transferred to the USEwindow by depressing the transfer button.Both the COMM and NAV, and USE andSTANDBY frequencies are stored in a NVM onpower down, and will be displayed when theunit is turned on. If an invalid frequency isdetected in the memory when power isappl ied, the COMM USE and STANDBYwindows will display 120.000 and the NAVUSE and STANDBY windows will display110.00.

In addition, when the smaller NAV frequencyselector knob is pulled on the KX-165, theradial of the active VOR is displayed in theSTANDBY window.

The KY196A is a VHF COMM transceiverwhich operates within the frequency range of118.000 MHz to 136.975 MHz in 25 kHzincrements (720 channels). The KY196Adisplay presents a USE and STANDBYfrequency with a T appearing between thefrequency numbers to indicate operation inthe transmit mode. The desired frequency isentered into the STANDBY window andtransferred to the use window by depressing

the transfer button. The COMM USE andSTANDBY frequencies are stored in a NVM onpower down, and will be displayed when theunit is turned on.

1-36. THREE OR FIVE PLACEINTERCOMMUNICATIONSYSTEM

Optional intercommunication and audiodistribution is accomplished by one KMA24H-71 audio control panel (Figure 1-4).

The KMA 24H-71 has separate speaker andheadphone isolation amplifiers which arepowered by separate circuit breakers on theavionic bus, providing a high degree of audiointegrity. The system is equipped with primaryand secondary headphone amplifiers servingthe pilot and the crew/intercom stationsrespectively. This provides for the pilot and/orcopilot to isolate from the intercom system.Headphone outputs contain the audioselected by the PHONE pushbuttons and theMIC rotary selector switch.

The system requires headsets or helmets witha 300 ohm impedance rating. Military typehelmets with an 8 ohm impedance rating arenot compatible and may cause damage to theaudio control panel.

The KMA 24H-71 audio panel MIC rotaryselector switch has seven positions which arenormally connected as follows:

POSITION FUNCTIONEMG Emergency1 VHF No. 12 VHF No. 23 Not Used4 Not Used5 Not UsedPA Aft Cabin Speaker


31 JAN 2007—Rev. 2———1-103

The KMA 24H-71 audio panel mixing pushbutton switches are normally connected asfollows:

The ICS mode rotary selector switch, locatedon the center console, has three modepositions: NORMAL, ISOLATE, and PRIVATE.Operation in the three modes is as follows.

1-36-A. NORMAL MODE

With the ICS mode rotary selector switchpositioned to NORMAL, and the KMA 24H-71audio panel MIC rotary selector switchposi t ioned to COMM 1 o r COMM 2 ( i fapplicable), all headphones will receive audiofrom all ICS inputs and selected audio controlpanel mixing push button selections.

To operate in the NORMAL (keyed ICS) mode,the VOX keying knob on the KMA 24H-71audio panel is to be turned fully counterclockwise. This wi l l a l low audio to a l lheadsets when the pilot or copilot cyclicswitch is keyed to the first position, thecopilot foot switch is keyed, or any of the aftICS drop chord switches are keyed.

To operate in the NORMAL (Hot MIC) mode,the VOX keying knob on the audio panel is tobe turned fully clockwise. This will allow

audio to all headsets by voice activation fromany position.

1-36-B. ISOLATE MODE

With the ICS mode rotary selector switchpositioned to ISOLATE, and the KMA 24H-71audio panel MIC rotary selector switchposi t ioned to COMM 1 o r COMM 2 ( i fapplicable), the pilot will be isolated from theintercom. Communication between the copilotand aft ICS stations will be available andaudio from the pilot shall not be heard in thecopilot, aft cabin speaker, or aft ICS headsets.Audio control panel mixing push buttonselections will only be heard through the pilotheadset.

To operate in the ISOLATE (keyed ICS) mode,the VOX keying knob on the KMA 24H-71audio panel is to be turned fully counterclockwise. This will allow audio between thecopilot and aft ICS positions when the copilotcyclic switch is keyed to the first position, thecopilot foot switch is keyed, or any of the aftICS drop chord switches are keyed.

To operate in the ISOLATE (Hot MIC) mode,the VOX keying knob on the audio panel is tobe turned fully clockwise. This will allowaudio between the copi lot and aft ICSpositions by voice activation.

1-36-C. PRIVATE MODE

With the ICS mode rotary selector switchpositioned to PRIVATE, and the KMA 24H-71audio panel MIC rotary selector switchposi t ioned to COMM 1 o r COMM 2 ( i fapplicable), the pilot and copilot will beisolated from the aft ICS stations.

Hot M IC (vo ice act ivated) aud io isautomatically provided to all positions in thePRIVATE mode regardless of the VOX keyingknob position. In addition, audio from theaudio control panel mixing push buttonselections will only be heard by the pilot andcopilot.

SWITCH FUNCTIONCOMM 1 VHF No. 1COMM 2 VHF No. 2COMM 3 Not UsedCOMM 4 Not UsedCOMM 5 Not UsedNAV 1 VOR/ILSNAV 2 Not UsedDME Not UsedMKR Not Used ADF ADFSPKR AUTO Audio from SPKR

PULL OUT switch


1-104———Rev. 2—31 JAN 2007

1-36-D. AFT CABIN MODE

Aft cabin speaker audio may be selected bypositioning the KMA 24H-71 audio panel MICrotary selector switch to PA. Keying the pilotor copilot cyclic stick switch to the secondposition (radio) will allow audio to be heard inthe aft cabin speaker. This will occur with theICS mode rotary selector switch positioned toNORMAL, ISOLATE, or PRIVATE.

1-37. EMERGENCYCOMMUNICATION SYSTEM

In the event of an intercommunication systemfa i lure , emergency communicat ion isavailable by positioning the KMA 24H-71audio panel MIC rotary selector switch toEMG. With VHF COMM 1 operating andselected to the required frequency, keying thepilots cyclic stick switch to the secondposi t ion ( rad io ) w i l l a l low two-waycommunication. Audio will only be availablein the pilots headset.

Additionally, the emergency communicationsystem can be activated with the ICS moderotary selector switch positioned to NORMAL,ISOLATE, or PRIVATE.

1-38. AVIONICS MASTER SWITCH

The AVIONICS MASTER switch a l lowsactivation and deactivation of all avionicscomponents simultaneously. Circuit breakersthat are connected to the avionics masterswitch are identified by triangles next to thecircuit breakers on the overhead console. Themain 28 VDC bus powers COMM 1 (VHF 1NAV/COMM) and the audio panel ( ICSPHONE).

1-39. MISCELLANEOUSINSTRUMENTS

Misce l laneous inst ruments inc lude acombined outside air temperature, clock, andvoltmeter indicator, and an hourmeter.

1-39-A. OUTSIDE AIR TEMPERATUREINDICATOR

The outside air temperature is included in amultifunction indicator mounted in the upperleft area of the instrument panel. The indicatoralso displays voltmeter and clock functions.

A button located on the center top of theinstrument changes the top display betweenOAT (Celsius and Fahrenheit) and volts (e.g.28E). When power is applied to the instrumentthe display defaults to the voltmeter reading.Pushing the button will change the display totemperature.

The temperature is taken by a probe mountedoutside the helicopter in the lower nosesection.

When all electrical power is turned off, theoutside air temperature display disappears.

1-39-B. HOURMETER

An hourmeter is located on the aft wall of thebattery compartment. It can only be viewedfrom the inside of the battery compartment.The hourmeter is a digital instrument thatregisters cumulative time, in hours andtenths. The hourmeter is powered by thehourmeter circuit breaker located in thebattery compartment. For the hourmeter torecord time, the hourmeter circuit breakermust be in, the engine NG must be greaterthan 55%, and the helicopter weight must beoff the landing gear (the helicopter must bein-flight).

1-39-C. VENTILATION AND DEFOGSYSTEM

A vent and defog system is installed at eachcrew station. Each side consists of a plenumwith a vent door, electric blower, windshielddefog nozzle, and a control cable.

Control cables, with knobs, are installedbelow either side of the instrument panel.Pulling the cable opens the exterior vent doorto allow outside ram air to enter the cabin


31 JAN 2007—Rev. 2———1-105

through an opening at the bottom of theplenum. The control cables lock in anyposition and are released by pressing thecenter button on each knob.

An electrically-driven axial flow blower ineach system provides airflow for ventilationand defogging when the helicopter is on theground or hovering. The blower intake takesair from the cabin and blows it through thewindshield defog nozzle onto the windshield.Both blowers are controlled by one DEFOGswitch located in the overhead console.

1-39-D. LIGHTING SYSTEM

The lighting systems include interior andexterior lighting. The interior lighting systemincludes a cockpit utility light, instrumentpanel and associated lighting, and aft cabinlighting. The exterior lighting system includeslanding, position, and anticollision lighting.

1-39-E. COCKPIT UTILITY LIGHT

A removable utility light is secured in abracket on the forward side of the controltube tunnel between the cockpit seat backs. Along , sp i ra l wound cord permi ts useanywhere in the cockpit.

The light will provide a blue or white lightdepending on the setting of the selectionswitch. It can be used as a spot or flood lightand the intensity of the light is controlled withthe BRT/DIM control knob. Power to thecockpit utility light is provided through a5-amp CKPT LIGHTS circuit breaker.

1-39-F. INSTRUMENT PANEL ANDASSOCIATED LIGHTING

The instrument lights are powered by 28 or 5VDC. An INSTR LIGHTS circuit breakerprov ides 28 VDC d i rect ly to cer ta ininstruments and also to a 5 volt power supply.Some instruments (primarily propulsioninstruments) are lighted by the 5 volt powersupply. Other instruments (primarily flightinstruments) are powered by 28 VDC. All

instrument lights are controlled by the INSTRLT rheostat located in the overhead console.The caution panel is supplied with 28V inbright and 15V in dim.

The caution/warning panel dimming isrestricted to one position 15V dimming. Todim the caution panel, the INSTR LT rheostatmust be positioned between dim and bright.Momentarily positioning the CAUT LT dimswitch to DIM will dim the caution panel lightsto a fixed dim mode.

Caution/warning panel segments FLOATTEST, BATTERY HOT, ENGINE OVERSPEED,ENGINE OUT, and RPM are not dimmable. Totest all of the caution panel segment lamps,press the CAUTION LT TEST switch. This willalso test the lamps for the FADEC modeswitch and NAV/GPS, and QUIET MODEswitch lamps (if installed). The PEDAL STOPswitch must be pressed to test its internallamps.

1-39-G. AFT CABIN LIGHTING

The aft cabin lighting system consists of twoindividual reading lights. Power is provided tothe lights from a 5-amp CKPT LIGHTS circuitbreaker and controlled through the CABIN/PASS LT switch and two individual readinglight switches.

With the CABIN/PASS LT switch positioned toOFF, all cabin lights will be OFF regardless ofthe position of the two reading light switches.With the CABIN/PASS LT switch positioned toCABIN LT, a l l cab in l igh ts w i l l be ONregardless of the position of the two readinglight switches. Positioning the CABIN/PASSLT switch to PASS LT will allow control of thereading lights through the reading lightswitches.

1-39-H. LANDING LIGHTS

The landing lights consist of one forward andone downward facing lamp. Both lamps areexposed for improved cooling.


1-106———Rev. 2—31 JAN 2007

Power is provided to the landing lightsthrough separate relays which are controlledby the LDG LIGHTS switch located on thecollective switch box. A 2-amp LDG LT CONTcircuit breaker is used to power the coils ofeach control relay and a 25-amp LDG LTPOWER circuit breaker is used to power thelanding lights through the control relays.

Positioning the LDG LIGHTS switch to FWDwill turn on the forward landing light andpositioning the switch to both will turn onboth the forward and downward landinglights.

1-39-I. POSITION LIGHTS

The position lights consist of a green lightlocated on the horizontal stabilizer rightvertical fin, a red light on the horizontalstabilizer left vertical fin, and a white lightlocated on the tail. Position lights are alsolocated on the lower cabin with a green lighton the right side, and a red light on the leftside.

Power is provided to the position lightsthrough a 5-amp POS LT circuit breakerswitch located on the overhead panel.

1-39-J. ANTICOLLISION LIGHT

The anticollision light consists of a single redstrobe light mounted on top of the vertical fin.Power is provided to the strobe from aseparate power supply, which is controlled bya 5-amp ANTI COLL LT circuit breaker switchlocated on the overhead panel.

1-40. EMERGENCY EQUIPMENT

Emergency equipment includes a portable fireextinguisher, a first aid kit, and optionalPointer 4000 ELT kit.

1-41. PORTABLE FIREEXTINGUISHER

A portable fire extinguisher is mountedbetween the cockpit seat backs.

1-42. FIRST AID KIT

The f i rs t a id k i t is suppl ied as looseequipment.

1-43. POINTER 4000 ELT

The Pointer 4000 ELT installation includes a4000-10 transmitter installed on the left side ofthe pedestal, a remote switch on the right sideof glare shield, and an external antennamounted on the forward upper cowl.

The Poin ter ELT is a se l f con ta inedemergency transmitter capable of manual orautomatic operat ion. I t is designed towithstand fo rced land ing and crashenvironment conditions. Automatic activationis accomplished by a deceleration sensinginertia switch. The inertia switch is designedto activate when the unit senses longitudinalinertia forces, as required in TSO-C91A.

To configure the ELT to automatically activatewith the remote switch installed, the Masterswitch on the ELT must be set to AUTO andthe remote switch must be set to AUTO. Thiswill allow the ELT to activate when the inertiaswitch senses predetermined decelerationlevel.

To override the inertia switch and turn on theELT manually, the remote switch must bepositioned to ON. This procedure may beused during unit testing or if an emergencysituation is imminent and pilot wishes toactivate ELT prior to emergency.

The RESET position of the remote switch isused to deact ivate and rearm the ELTtransmitter to AUTO mode after automaticactivation by the inertia switch. Helicopterpower is required for remote reset. In case ofinadvertent activation of ELT transmitter withhelicopter power OFF, turn helicopter powerON, position remote switch to RESET, andthen back to AUTO.

After a forced landing or helicopter accident,if helicopter receiver is operable, listen on


121.5 MHz for ELT transmissions. The rangeof ELT varies according to weather andtopography. In general, the swept tone signalcan be heard up to 30 miles by a searchhelicopter at 10,000 feet. It is recommended tostay close to the helicopter to permit easierspotting by airborne searches.

I t may a lso be des i rab le to use ELTtransmitter in the portable mode due to abroken or disabled whip antenna, severedantenna coax cable , danger of f i re orexplos ion , temperatu re ext remes inhelicopter, poor transmitting location, orwater ditching with forced evacuation.

To remove transmitter from helicopter:

1. Bend switch guard away from unitMaster Switch and place switch inOFF position.

2. Disconnect remote antenna coaxcable.

3. Disconnect remote switch cable.

4. Remove telescopic antenna fromstowage clips. Unlatch ELTtransmitter hold down strap andremove unit from bracket.

5. Insert telescopic antenna into ANTreceptacle. Extend antenna fully.

6. Turn ELT transmitter Master switchto ON position. DO NOT USE AUTOPOSITION.

Consider factors such as terrain, temperature,and precipitation when choosing a locationfor the ELT transmitter to radiate from. Thebest transmission may be obtained bykeeping the antenna vertical and settingtransmitter upright on a metallic surface. Ifterrain prohibits good transmission (such asdeep valley or canyon), place the transmitteron high ground or hold in hand in high place.

As cold temperatures have a direct effect onELT transmitter battery life, the life of thebattery pack can be extended by placing theELT transmitter inside a jacket or coat to keepthe battery warm. Let antenna extend outsidejacket. Keep all moisture and ice away fromthe antenna connection and the remoteconnector pins.

Ensure ELT t ransmi t te r is turned ONcontinuously (day and night), until rescueteam appears.

For additional information, refer to the PointerAircraft Emergency Locator TransmitterOperation and Installation booklet.

31 JAN 2007—Rev. 2———1-107/1-108


31 JAN 2007—Rev. 2———2-1

Section 2HANDLING AND SERVICING

2TABLE OF CONTENTS

Paragraph PageSubject Number NumberGround Handling.................................................................................... 2-1 ........... 2-3Covers and Tie-downs........................................................................... 2-2 ........... 2-3

Cover — Engine Inlet......................................................................... 2-2-A ....... 2-3Cover — Pitot Tube............................................................................ 2-2-B ....... 2-3Cover — Engine Exhaust .................................................................. 2-2-C ....... 2-3Cover — Oil Cooler Blower Inlet Duct.............................................. 2-2-D ....... 2-3Tie-down — Main Rotor..................................................................... 2-2-E ....... 2-5Tie-down — Tail Rotor....................................................................... 2-2-F........ 2-5Parking — Normal and Turbulent Conditions (Winds Up to50 Knots)............................................................................................. 2-2-G....... 2-6Mooring (Winds Above 50 Knots) .................................................... 2-2-H ....... 2-6

Fuels........................................................................................................ 2-3 ........... 2-7 Fuel System Servicing...................................................................... 2-3-A ....... 2-7

Oils .......................................................................................................... 2-4 ........... 2-8Engine Oils ......................................................................................... 2-4-A ....... 2-8Engine Oil System Servicing ............................................................ 2-4-B ....... 2-8Transmission and Tail Rotor Gearbox Oils ..................................... 2-4-C ....... 2-9Transmission and Tail Rotor Gearbox Servicing............................ 2-4-D ....... 2-9

Oil Change — Different Specification .......................................... 2-4-D-1 .... 2-10Deleted ............................................................................................ 2-4-D-2 .... 2-10

Hydraulic Fluids ..................................................................................... 2-5 ........... 2-10Hydraulic System Servicing.............................................................. 2-5-A ....... 2-10

FIGURES


Covers and Tie-downs........................................................................... 2-1 ........... 2-4


2-2———Rev. 2—31 JAN 2007

TABLES


Commercial Fuels — ASTM D-1655 (Type A and A-1)........................ 2-1........... 2-11Commercial Fuels — ASTM D-6615 (Type B)...................................... 2-2........... 2-12Military Fuels.......................................................................................... 2-3........... 2-13Engine Oils ............................................................................................. 2-4........... 2-14Transmission and Tail Rotor Gearbox Oils......................................... 2-5........... 2-16Hydraulic Fluids — MIL-H-5606 (NATO H-515).................................... 2-6........... 2-17


31 JAN 2007—Rev. 2———2-3

Section 2HANDLING AND SERVICING

22-1. GROUND HANDLING

Ground handling of the helicopter consists oftowing, parking, securing, and mooring.Model 205 or 206 ground handling wheels areused for towing. Refer to the BHT-407-MM-2,Chapter 9 for more detailed ground handlinginformation.

CAUTION

DO NOT TOW THE HELICOPTER IFTHE GROSS WEIGHT IS MORE THAN5000 POUNDS (2270 KG) FORMODEL 205 GROUND HANDLINGWHEELS, OR 4450 POUNDS (2020KG) FOR MODEL 206 GROUNDHANDLING WHEELS.

2-2. COVERS AND TIE-DOWNS

Protective covers and tie-downs are furnishedas loose equipment and are used for parkingand mooring of helicopter (Figure 2-1).Additional equipment such as ropes, cables,clevises, ramp tie-downs, or dead mantie-downs will be required during mooring.

2-2-A. COVER — ENGINE INLET

Engine inlet plug assemblies are red andflame resistant, and each cover is attachedwith a red streamer stenciled in white letters,REMOVE BEFORE FLIGHT. To install theengine inlet plug, make sure that the side

marked TOP is facing upwards. Push theengine inlet plug into the engine air inlet.

2-2-B. COVER — PITOT TUBE

WARNING

THE PITOT TUBE CAN BE HOT.

Pitot tube cover assembly is red, flameresistant, and attached with a red streamerstenciled in white letters, REMOVE BEFOREFLIGHT. To install pitot tube cover, push itover the pitot tube. Attach the cord.

2-2-C. COVER — ENGINE EXHAUST

Engine exhaust cover is red and flameresistant, and includes a red streamerstenciled in white letters, REMOVE BEFOREFLIGHT. A 1/4 inch diameter elastic tie-cord isattached to cover for securing to engineexhaust. To install the engine exhaust cover,push it over the exhaust tailpipe. Attach thetie-cord.

2-2-D. COVER — OIL COOLER BLOWERINLET DUCT

Oil cooler blower inlet duct plug assembliesare red and flame resistant, and each cover isattached with a red streamer stenciled inwhite letters, REMOVE BEFORE FLIGHT. Toinstall the inlet plug, push it into the oil coolerblower inlet duct.


2-4———Rev. 2—31 JAN 2007

Figure 2-1. Covers and Tie-downs407_MD-1_02_0001

DETAIL C

SEE DETAIL C

SEE DETAIL B

DETAIL B

DETAIL A

SEE DETAIL A

1.

2.

3.

4.

5.

6.

7.

8.

Pitot tube cover assembly

Engine inlet plug assembly

Engine exhaust cover

Tie-down assembly

Sock assembly

Line assembly

Tail rotor tie-down strap

Oil cooler blower inlet duct plugs

50.0 IN.(1270.0 mm)

60.0 LB (27.2 kg)

MAXIMUM LOAD

60.0 LB (27.2 kg)

MAXIMUM LOAD

MAXIMUM

DEFLECTION

6

4

5

DETAIL D

SEE DETAIL D

7

6

1

7

4

5

3

68

2


31 JAN 2007—Rev. 2———2-5

2-2-E. TIE-DOWN — MAIN ROTOR

For each main rotor blade, there is a mainrotor tie-down assembly. Each tie-downassembly has a sock assembly and a lineassembly. Use these assemblies to attach theblades to the landing gear crosstubes. Thesock assembly is red and has a red streamerattached to it stenciled with white letters:REMOVE BEFORE FLIGHT. The line assemblyis made of 0.19 inch (4.83 mm) diameter nylonand has a ring and an attached flag. The flagis stenciled with the letters FWD BLADES orAFT BLADES.

Install the main rotor tie-down assemblies asfollows:

CAUTION

DO NOT CAUSE THE MAIN ROTORBLADES TO BEND MORE THAN THELIMITS SHOWN IN FIGURE 2-1,DETAIL A.

NOTE

At the same time that you align themain rotor blades, align the tail rotorblades with the vertical fin. This willmake it possible to install the tailrotor tie-down.

1. Turn the main rotor blades until thereare two blades aft of the fuselagestation of the main rotor hub. Whenyou look down at the helicopter, thefour blades make an X over thevertical center line of the fuselage.

2. Install the two FWD BLADES sockassemblies on the ends of the mainrotor blades that are forward of thefuselage station of the main rotorhub.

3. Put a line assembly around eachoutboard end of the forwardcrosstube of the landing gear.

NOTE

Rings are pre-set to apply thenecessary tension to the forward andaft main rotor blades.

4. Attach the snaps of the two lineassemblies to the rings of the twoFWD BLADES sock assemblies.

5. Install the two AFT BLADES sockassemblies on the ends of the two aftmain rotor blades.

6. Put a line assembly around eachoutboard end of the crosstube of thelanding gear.

7. Attach the snaps of the two lineassemblies to the rings of the twoAFT BLADES sock assemblies.

2-2-F. TIE-DOWN — TAIL ROTOR

The tail rotor tie-down strap is made of 0.025 x1.0 x 92.0 inches (0.635 x 25 x 2340 mm) nylonwebbing. It is red and stenciled with whiteletters REMOVE BEFORE FLIGHT.

Install the tail rotor tie-down as follows:

CAUTION

DO NOT TIE DOWN TAIL ROTOR TOEXTENT THAT TAIL ROTOR BLADEFLEXES. THE APPLIED FORCE OFTHE TIE-DOWN SHOULD PROVIDE ALIGHT CONTACT BETWEEN THETAIL ROTOR YOKE AND THEFLAPPING STOP.

1. Turn the main rotor blades until thereare two blades aft of the fuselagestation of the main rotor hub. Whenyou look down at the helicopter, thefour blades should make an X overthe vertical center line of thefuselage. Align the tail rotor with thevertical fin.


2-6———Rev. 2—31 JAN 2007

2. Leaving enough strap material toextend around the vertical fin, wrapthe tail rotor tie-down strap aroundthe upper tail rotor blade once(Figure 2-1, Detail D).

3. Position the loose ends of the straparound the upper half of the verticalfin and tie the two ends together.

2-2-G. PARKING — NORMAL ANDTURBULENT CONDIT IONS(WINDS UP TO 50 KNOTS)

When winds are forecast to be light or up to50 knots, park helicopter pointed in directionfrom which you expect the highest winds.

Moor helicopter as follows:

1. Hover, taxi, or tow helicopter to thespecified parking area.

2. Remove ground handling gear (ifinstalled).

3. Attach the main and tail rotor bladetie-downs.

4. Install the engine air inlet plugs, oilcooler blower inlet duct plugs, pitottube, and the engine exhaust covers.

5. Tighten the friction locks on the flightcontrols.

6. Make sure that all switches are in theOFF position.

7. Disconnect the battery.

8. Close and safety all of the doors,windows, cowlings, and accesspanels.

9. If helicopter is parked outside in aheavy dew environment, purgelubricate all of the control bearingsthat are open to the air. Do this onceevery 7 days. Make sure no voidsexist that could trap moisture.

2-2-H. MOORING (WINDS ABOVE 50KNOTS)

When winds above 50 knots are forecast, parkhelicopter pointed in direction from which youexpect the highest winds.

Moor helicopter as follows:

CAUTION

WHEN WINDS ABOVE 75 KNOTSARE FORECAST, PUT THEHELICOPTER IN A HANGAR ORMOVE IT TO AN AREA WHERE ITWILL NOT BE AFFECTED BY THEWEATHER. FLYING OBJECTSDURING HIGH WINDS CAN CAUSEDAMAGE TO THE HELICOPTER.

NOTE

If the correct ramp tie-downs are notavailable, park the helicopter on anunpaved area. Use the dead mantie-downs. Point the helicopter intothe wind and remove the groundhandling wheels.

1. Attach helicopter to the ramptie-downs.

NOTE

Use a mooring clevis at each of thethree jack fittings. This will let youuse a rope with a larger diameter.

2. Attach the cable, rope, or themanufactured tie-downs to thehelicopter jack fittings.

CAUTION

DO NOT CAUSE THE MAIN ROTORBLADES TO BEND MORE THAN 50.0INCHES (1270 MM), AS SHOWN INFIGURE 2-1.


30 APR 2008—Rev. 4———2-7

3. Attach the main and tail rotortie-downs.

4. If time and storage space areavailable, remove the main rotorblades and put them in a safebuilding.

NOTE

Put all of the red streamers inside anaccess door so that they will not flapin the wind.

5. Install the engine air inlet plugs, oilcooler blower inlet duct plugs, pitottube cover, and engine exhaustcover.

6. Tighten friction locks on the flightcontrols.

CAUTION

MAKE SURE THAT ALL OF THESWITCHES ARE IN THE OFFPOSITION, AND THAT ALL OF THECIRCUIT BREAKERS ARE OPEN.

7. Disconnect the battery.

8. Close and safety all of the doors,windows and access panels.

9. Refuel the helicopter to its maximumcapacity.

CAUTION

SAFETY OR REMOVE ALL OF THEEQUIPMENT AND OBJECTS IN THEAREA. OTHERWISE, THE WIND CANBLOW THE OBJECTS AGAINST THEHELICOPTER AND CAUSE DAMAGE.

10. Safety or remove all of the equipmentand objects in the area.

11. When the winds stop, examine thehelicopter for damage.

2-3. FUELS

Fuels conforming to following the commercialand military specifications are approved:

Refer to the BHT-407-FM-1 for fuel limitationsand use of anti-icing additive.

Fuel listings (Table 2-1 through Table 2-3) areprovided for convenience of operator. It shallbe the responsibility of the operator and thefuel supplier to ensure fuel used in helicopterconforms to one of approved specifications.

Re fer to Rol ls -Royce Opera t ion andMaintenance Manua l fo r a l te rna te oremergency fuels.

2-3-A. FUEL SYSTEM SERVICING

Fuel system contains two interconnectedcells that are serviced through a single fuelport located on right side of helicopter. Agrounding jack is provided near fueling port.An electric sump drain is located in both the

SPECIFICATION OAT RANGE

ASTM D-1655, Jet A or A-1

Above -32°C(-25°F)

ASTM D-6615, Jet B Any OAT

MIL-DTL-5624,Grade JP-4 (NATO F-40)

Any OAT


Above -32°C(-25°F)


Above -32°C(-25°F)

Total capacity 130.5 U.S. gallons (493.9 L)

Usable fuel 127.8 U.S. gallons (483.7 L)

Unusable fuel 2.7 U.S. gallons (10.2 L)

Undrainable fuel (included in unusable fuel)

0.7 U.S. gallon(2.6 L)


2-8———Rev. 4—30 APR 2008

forward and aft tanks. They are activated bybuttons located on the right aft lower side offuselage. Battery switch must be ON (orexternal power applied) and fuel valve switchmust be OFF to activate sump drains.

2-4. OILS

Approved oils and vendors are listed in thissection for convenience of operator.

An appropriate entry shal l be made inhelicopter logbook when oil has been addedto engine, transmission, or tail rotor gearbox.Entry shall show specification and brandname of oil used to prevent inadvertentmixing of oils.

CAUTION

DO NOT MIX OILS OF DIFFERENTSPECIFICATIONS. IF OILS BECOMEMIXED, SYSTEM SHALL BEDRAINED, FLUSHED, ANDREFILLED WITH PROPERSPECIFICATION OIL.

2-4-A. ENGINE OILS

Cer ta in o i ls con forming to fo l lowingspecifications are approved for use in engine:

NOTE

As per Rolls-Royce, the preferredengine oils for MIL-PRF-23699 areMobil Jet Oil 254 and Aeroshell 560.

Engine oils (Table 2-4) shall meet enginemanufac turer ' s approva l . Consu l tRolls-Royce Operation and MaintenanceManual for use of oil brands not listed herein.

NOTE

Because of availability, reducedcoking, and bet ter lubr icat ingqualities at higher temperatures,qualified MIL-PRF-23699 oils arepreferred by engine manufacturer.

NOTE

Long term use of DOD-PRF-85734 oilmay increase probability of sealleakage in accessory gearbox.

Refer to the BHT-407-FM-1 for engine oillimitations.

2-4-B. ENGINE OIL SYSTEM SERVICING

Capacity: 6.0 U.S. quarts (5.7 L).

Engine oil tank is located under aft fairing,and access doors are provided for filling anddraining oil tank. A sight glass and filler capdip stick are provided to determine quantity ofoil in tank.

NOTE

If helicopter engine has been shutdown for more than 15 minutes,scavenge oil could have drained intogearbox. Dry motor run engine for 30seconds before checking oil level. Ifnot accomplished, a false highengine o i l consumpt ion ra teindication or overfilling of oil tankcould result. Do not overfill engine oiltank.

NOTE

MIL-PRF-23699 and DOD-PRF-85734oils are not approved for use inambient temperatures below -40°C(-40°F). When changing to an oil of a


MIL-PRF-7808(NATO O-148)

Any OAT


OAT above -40°C(-40°F)

DOD-PRF-85734———



30 APR 2008—Rev. 4———2-9

different specification, system shallbe drained and flushed.

Re fer to Rol ls -Royce Opera t ion andMaintenance Manua l for serv ic inginstructions and oil filter change procedures.

2-4-C. TRANSMISSION AND TAILROTOR GEARBOX OILS

Oils conforming to following specificationsare approved for use in transmission and tailrotor gearbox (Table 2-5):

NOTE

I t is recommended thatDOD-PRF-85734 oi l be used intransmission and tail rotor gearboxto maximum extent al lowed bytemperature limitations. Refer to theBHT-407-FM-1 for transmission andtail rotor gearbox oil limitations.

2-4-D. TRANSMISSION AND TAILROTOR GEARBOX SERVICING

Sight glasses are provided to determinequantity of oil in transmission and tail rotorgearbox.

NOTE

Transmission oil may partially draininto freewheel assembly after shutdown. When checking oil levels,consider this and slope of helicopterlanding surface. If not considered, afalse oil quantity indication oroverfilling of gearbox could result.

NOTE

DOD-PRF-85734 oil is not approvedfor use in ambient temperaturesbelow -40°C (-40°F). When changingto an oil of a different specification,system shall be drained and flushed.

When adding oil to the main transmission ortail rotor gearbox, the identical brand andspecification of oil already in each gearboxshall be used. However, in circumstanceswhere emergency top-off or inadvertentmixing may occur, it is acceptable to use oilwith a different brand name within the samespecification. No further action will berequired until the next scheduled oil change,provided there is no indication of foggy orhazy oil appearance in the sight gauge.

If oils of different specifications have beenmixed or a foggy or hazy oil appearanceexists, accomplish the required steps perparagraph 2-4-D-1.

Refer to the BHT-407-MM-2, Chapter 12 fordetailed procedures for draining oil andchanging filters.

2-4-D-1. OIL CHANGE — DIFFERENTSPECIFICATION

When changing to an oil of a dif ferentspecification, accomplish following steps:

NOTE

Refer to the BHT-407-MM-2, Chapter12 for the maintenance instructions.

1. Drain transmission, freewheel unit,and tail rotor gearbox.

2. Replace transmission oil filter.

3. Service transmission with properamount of approved oil.


DOD-PRF-85734—



Below -18°C(0°F)

Transmission capacity 5.0 U.S. quarts(4.7 L)

Tail rotor gearbox capacity

0.33 U.S. quarts(0.31 L)


2-10———Rev. 4—30 APR 2008

4. Service tail rotor gearbox with properamount of approved oil.

5. Operate helicopter for not less than30 minutes nor longer than 5 hours.

6. Drain transmission, freewheel unit,and tail rotor gearbox.

7. Service transmission with properamount of approved oil.

8. Service tail rotor gearbox with properamount of approved oil.

9. During first 100 hours of operationwith new oil, check oil sight glassesclosely for indications of foggy orhazy appearance. If these indicationsoccur, repeat step 6 through step 9until eliminated.

2-5. HYDRAULIC FLUIDS

Hydraulic fluids listed in Table 2-6 conform toMIL-PRF-5606 (NATO H-515) and are approved

for use in hydraulic flight control system androtor brake.

2-5-A. HYDRAULIC SYSTEM SERVICING

Hydraulic reservoir is located on top offuselage, forward of transmission, and underforward fairing. A sight glass is provided todetermine quantity of hydraulic fluid inreservoir.

Service hydraulic system as follows:

1. Open and support top of forwardfairing.

2. Remove cap and fill reservoir untilsight glass is full of hydraulic fluid.

3. Secure cap and fairing.

Reservoir capacity 1.0 U.S. pint(0.5 L)


30 APR 2008—Rev. 4———2-11

Table 2-1: Commercial Fuels — ASTM D-1655 (Jet A and A-1)

FUEL VENDOR ASTM D-1655, JET APRODUCT NAME

ASTM D-1655, JET A-1PRODUCT NAME

American Oil and Supply American Jet Fuel Type A American Jet Fuel Type A-1

ARCO (Atlantic Richfield) Arcojet A Arcojet A-1

Boron Oil Jet A Kerosene Jet A-1 Kerosene

British-American B-A Jet Fuel JP-1

British Petroleum B.P. Jet A B.P. A.T.K.

California-Texas Caltex Jet A-1

Chevron Chevron Jet A-50 Chevron Jet A-1

Cities Service Citgo Turbine Type A

Continental Conoco Jet-50 Conoco Jet-60

Exxon Co. USA Exxon Turbo Fuel A Exxon Turbo Fuel A-1

Exxon International Esso Turbo Fuel A-1

Gulf Oil Gulf Jet A Gulf Jet A-1

Mobil Oil Mobil Jet A Mobil Jet A-1

Phillips Petroleum Philjet A-50

Pure Oil Purejet Turbine Fuel Type A Purejet Turbine Fuel Type A-1

Shell Oil AeroShell Turbine Fuel 640 AeroShell Turbine Fuel 650

Standard Oil of BritishColumbia

Chevron Jet Fuel A-50 Chevron Jet Fuel A-1

Standard Oil of California Chevron Jet Fuel A-50 Chevron Jet Fuel A-1

Standard Oil of Indiana American Jet Fuel Type A American Jet Fuel Type A-1

Standard Oil of Kentucky Standard Turbine Fuel A-50 Standard Turbine Fuel A-1

Standard Oil of New Jersey Standard Jet A Standard Jet A-1

Standard Oil of Ohio Jet A Kerosene Jet A-1 Kerosene

Standard Oil of Texas Chevron Avjet A Chevron Avjet A-1

Union Oil 76 Turbine Fuel


2-12———Rev. 4—30 APR 2008

Table 2-2: Commercial Fuels — ASTM D-6615 (Jet B)

FUEL VENDOR ASTM D-6615PRODUCT NAME

American Oil and Supply American JP-4

ARCO (Atlantic Richfield) Arcojet B

British-American B-A Jet Fuel JP-4

British Petroleum B.P. A.T.G.

California-Texas Caltex Jet B

Chevron Chevron Jet B

Continental Conoco JP-4

Exxon Co. USA Exxon Turbo Fuel 4

Exxon International Esso Turbo Fuel 4

Gulf Oil Gulf Jet B

Mobil Oil Mobil Jet B

Phillips Petroleum Philjet JP-4

Shell Oil AeroShell Turbine Fuel JP-4

Standard Oil of California Chevron Jet Fuel B

Standard Oil of Indiana American JP-4

Standard Oil of Kentucky Standard Turbine Fuel B

Standard Oil of New Jersey Standard Jet B

Standard Oil of Texas Chevron Jet Fuel B

Texaco Texaco Avjet B

Union Oil Union JP-4


30 APR 2008—Rev. 4———2-13

Table 2-3: Military Fuels

COUNTRY NATO F-34(JP-8 TYPE)

NATO F-40(JP-4 TYPE)

NATO F-44(JP-5 TYPE)

Belgium BA-PF-7 BA-PF-2 3-GP-24

Canada 3-GP-22 3-GP-24

Denmark D. Eng. R.D. 2453 MIL-DTL-5624,Grade JP-4

France AIR 3405 AIR 3407 AIR 3404

Germany VTL-9130-006 VTL-9130-007 VTL-9130-010

Greece MIL-DTL-5624,Grade JP-4

Italy AA-M-C.141 AER-M-C.142 AA-M-C.143

Netherlands D. Eng. R.D. 2453 MIL-DTL-5624,Grade JP-4

D. Eng. R.D. 2498

Norway MIL-DTL-5624,Grade JP-4

Portugal AIR 3405 MIL-DTL-5624,Grade JP-4

Turkey MIL-DTL-5624,Grade JP-4

United Kingdom D. Eng. R.D. 2453 D. Eng. R.D. 2454 D. Eng. R.D. 2498D. Eng. R.D. 2452

United States MIL-DTL-83133,Grade JP-8

MIL-DTL-5624,Grade JP-4

MIL-DTL-5624,Grade JP-5


2-14———Rev. 4—30 APR 2008

Table 2-4: Engine Oils

VENDOR PRODUCT NAME

SPECIFICATION MIL-PRF-7808 (NATO O-148) (FOR OAT ABOVE -40°C/-40°F)

Air BP BP Turbo Oil 2389

American Oil and Supply American PQ Lubricant 6899

Bray Oil Brayco 880H

Mobil Oil Mobil Avrex S Turbo 256Mobil RM-184AMobil RM-201A

Stauffer Chemical Stauffer Jet I

SPECIFICATION MIL-PRF-23699 (NATO O-156) OILS (FOR OAT ABOVE -40°C/-40°F)


American Oil and Supply American PQ Lubricant 6700

Caltex Petroleum Caltex RPM Jet Engine Oil 5

Castrol Brayco 899GCastrol 205

Chevron International Chevron Jet Engine Oil 5

Hatco Chemical Hatcol 3211

Mobil Oil Mobil Jet Oil IIMobil Jet Oil 254

Royal Lubricants Royco Turbine Oil 500Royco Turbine Oil 560

Shell Oil AeroShell Turbine Oil 500AeroShell Turbine Oil 560

Stauffer Chemical Stauffer Jet II (6924)

SPECIFICATION DOD-PRF-85734 (FOR OAT ABOVE -40°C/-40°F)


Royal Lubricants Royco Turbine Oil 555

Shell International Aeroshell Turbine Oil 555


30 APR 2008—Rev. 4———2-15

Table 2-5: Transmission and Tail Rotor Gearbox Oils

VENDOR PRODUCT NAME

SPECIFICATION MIL-PRF-7808 (NATO O-148) (FOR OAT BELOW -18°C/0°F)

Air BP BP Turbo Oil 2389BP Turbo Oil 2391

Burmah-Castrol (UK) Ltd. Castrol 399

Castrol Brayco 880Castrol 399

Hatco Chemical Hatcol 1278Hatcol 1280

Hexagon Enterprises Metrex AF Oil 01, 02, 07

Huls America AOSyn Jet IIIPQ Turbine Oil 4236PQ Turbine Oil 4706PQ Turbine Oil 4707PQ Turbine Oil 8365PQ Turbine Oil 9900

Mobil Oil RM-248ARM-272A

NYCO, S.A. Turbonycoil 160

Royal Lubricants Royco 808

Shell International AeroShell Turbine Oil 308

SPECIFICATION DOD-PRF-85734 (FOR OAT ABOVE -40°C/-40°F)


Royal Lubricants Royco Turbine Oil 555

Shell International Aeroshell Turbine Oil 555


2-16———Rev. 4—30 APR 2008

Table 2-6: Hydraulic Fluids — MIL-PRF-5606 (NATO H-515)

VENDOR PRODUCT NAME

Arpol Petroleum Arpolair 5606

Castrol Brayco Micronic 756

Castrol Canada Castrol Aero HF515

Chevron USA Chevron Aviation Hydraulic Fluid E(PED 5597)Chevron PED 6062Chevron PED 6063Chevron PED 6064Chevron PED 6065

Convoy Oil Convoy 606

Esso SAF Esso Fluid Aviation Invarol FJ13

Hexagon Enterprises Metrex Hydrol 1

Huls America PQ 4140PQ 9300PQ 9301PQ 9302PQ 9309PQ 9310PQ 9311

Mobil Oil Mobil Aero HFE

NYCO S.A. NYCO Hydraunycoil FH51

Rohm & Haas PA 4394

Royal Lubricants Royco 756

Shell International AeroShell 41

Technolube Technolube FB003


Section 3CO NVERSIO N CHARTS AND TABLES

3


Subject

TABLE OF CONTENTS

INTRODUCTION ......................................................................... 3-1.................. 3-3CONVERSION TABLES ............................................................. 3-2.................. 3-3

Title Number

TableNumberPage

LIST OF TABLES

Celsius to Fahrenheit conversion ............................................ 3-1.................. 3-4Gallons to liters conversion...................................................... 3-2.................. 3-5Inches to millimeters conversion ............................................. 3-3.................. 3-5Feet to meters conversion ........................................................ 3-4.................. 3-6Pounds to kilograms conversion ............................................. 3-5.................. 3-6Velocity conversion ................................................................... 3-6.................. 3-7Standard atmosphere ................................................................ 3-7.................. 3-8Barometric pressure conversion.............................................. 3-8.................. 3-9

09 Dec 2002 Reissue 3-1/3-2

g:/Bell/407_MD-1/407-md-3.pdf



09 Dec 2002 Reissue 3-3


Section 3CO NVERSIO N CHARTS AND TABLES

3

3-1. INTRODUCTION

This section contains additional informationwhich may be useful for operational planningbut which is not required for inclusion in theRotorcraft Flight Manual. Additional data maybe developed and included as appropriate.

3-2. CONVERSION TABLES

The conversion tables (Tables 3-1 through 3-8) provide useful information to assist in flightplanning and operations.


3-4 Reissue 09 Dec 2002

Table 3-1. Celsius to Fahrenheit conversion

CELSIUS TO FAHRENHEITCONVERSION TABLE

°F = (°C x 1.8) + 32°°C = (°F – 32°) x .555

C. →→→→ F. C. ←←←← F. C. →→→→ F. C. ←←←← F. C. →→→→ F. C. ←←←← F.

-62.2 -80 -112.0 71.1 160 320.0 260.0 500 932.0-56.7 -70 -94.0 73.9 165 329.0 265.6 510 950.0-51.1 -60 -76.0 76.7 170 338.0 271.1 520 968.0-45.6 -50 -58.0 79.4 175 347.0 276.7 530 986.0-40.0 -40 -40.0 82.2 180 356.0 282.2 540 1004.0-34.4 -30 -22.0 85.0 185 365.0 287.8 550 1022.0-31.7 -25 -13.0 87.8 190 374.0 293.3 560 1040.0-28.9 -20 -4.0 90.6 195 383.0 298.9 570 1058.0-26.1 -15 5.0 93.3 200 392.0 304.4 580 1076.0-23.3 -10 14.0 96.1 205 401.0 310.0 590 1094.0-20.6 -5 23.0 98.9 210 410.0 315.6 600 1112.0-17.8 0 32.0 101.7 215 419.0 326.7 620 1148.0-15.0 5 41.0 104.4 220 428.0 337.8 640 1184.0-12.2 10 50.0 107.2 225 437.0 348.9 660 1220.0

-9.4 15 59.0 110.0 230 446.0 360.0 680 1256.0-6.7 20 68.0 112.8 235 455.0 371.1 700 1292.0-3.9 25 77.0 115.6 240 464.0 382.2 720 1328.0-1.1 30 86.0 118.3 245 473.0 393.3 740 1364.01.6 35 95.0 121.1 250 482.0 404.4 760 1400.04.4 40 104.0 126.7 260 500.0 415.6 780 1436.07.2 45 113.0 132.2 270 518.0 426.7 800 1472.0

10.0 50 122.0 137.8 280 536.0 437.8 820 1508.012.8 55 131.0 143.3 290 554.0 454.4 850 1562.015.6 60 140.0 148.9 300 572.0 482.2 900 1652.018.3 65 149.0 154.4 310 590.0 510.0 950 1742.021.1 70 158.0 160.0 320 608.0 537.7 1000 1832.023.9 75 167.0 165.6 330 626.0 565.5 1050 1922.026.7 80 176.0 171.1 340 644.0 593.3 1100 2012.029.4 85 185.0 176.7 350 662.0 621.1 1150 2102.032.2 90 194.0 182.2 360 680.0 648.8 1200 2192.035.0 95 203.0 187.8 370 698.0 676.6 1250 2282.037.8 100 212.0 193.3 380 716.0 704.4 1300 2372.040.6 105 221.0 198.9 390 734.0 732.2 1350 2462.043.3 110 230.0 204.4 400 752.0 760.0 1400 2552.046.1 115 239.0 210.0 410 770.0 787.7 1450 2642.048.9 120 248.0 215.6 420 788.0 815.5 1500 2732.051.7 125 257.0 221.1 430 806.0 843.3 1550 2822.054.4 130 266.0 226.7 440 824.0 871.1 1600 2912.057.2 135 275.0 232.2 450 842.0 898.8 1650 3002.060.0 140 284.0 237.8 460 860.0 926.6 1700 3092.062.8 145 293.0 243.3 470 878.0 954.4 1750 3182.065.6 150 302.0 248.9 480 896.0 982.2 1800 3272.068.3 155 311.0 254.4 490 914.0 1010.0 1850 3362.0

(TABLE I.D. 910630)



Table 3-2. Gallons to liters conversion

GALLONS TO LITERS CONVERSION TABLE

U.S. GALLON

IMPERIALGALLON LITER

U.S.GALLON

IMPERIALGALLON LITER

10 8.33 37.85 170 141.55 643.4520 16.65 75.71 180 149.86 681.3030 24.98 113.56 190 158.20 719.1640 33.31 151.42 200 166.52 757.1850 41.63 189.27 210 174.84 795.0360 49.96 227.13 220 183.18 832.8970 58.28 264.98 230 191.50 870.7480 66.61 302.83 240 199.84 908.6090 74.94 340.69 250 208.14 946.45

100 83.26 378.54 260 216.48 984.45110 91.59 416.35 270 224.82 1022.16120 99.92 454.20 280 233.14 1060.01130 108.24 492.05 290 241.56 1097.87140 116.57 529.90 300 249.80 1135.62150 124.90 567.75 310 258.12 1173.47160 133.22 605.60 320 266.44 1211.33

(TABLE I.D. 910628)

Table 3-3. Inches to millimeters conversion

INCHES TO MILLIMETERSCONVERSION TABLE

Inches

0 1 2 3 4 5 6 7 8 9

mm mm mm mm mm mm mm mm mm mm

0 −−−− 25.4 50.8 76.2 101.6 127.0 152.4 177.8 203.2 228.610 254.0 279.4 304.8 330.2 355.6 381.0 406.4 431.8 457.2 482.620 508.0 533.4 558.8 584.2 609.6 635.0 660.4 685.8 711.2 736.630 762.0 787.4 812.8 838.2 863.6 889.0 914.4 939.8 965.2 990.640 1016.0 1041.4 1066.8 1092.2 1117.6 1143.0 1168.4 1193.8 1219.2 1244.650 1270.0 1295.4 1320.8 1346.2 1371.6 1397.0 1422.4 1447.8 1473.2 1498.660 1524.0 1549.4 1574.8 1600.2 1625.6 1651.0 1676.4 1701.8 1727.2 1752.670 1778.0 1803.4 1828.8 1854.2 1879.6 1905.0 1930.4 1955.8 1981.2 2006.680 2032.0 2057.4 2082.8 2108.2 2133.6 2159.0 2184.4 2209.8 2235.2 2260.690 2286.0 2311.4 2336.8 2362.2 2387.6 2413.0 2438.4 2463.8 2489.2 2514.6

100 2540.0 2565.4 2590.8 2616.2 2641.6 2667.0 2692.4 2717.8 2743.2 2768.6(TABLE I.D. 910627)



Table 3-4. Feet to meters conversion

FEET TO METERSCONVERSION TABLE

Feet

0 1 2 3 4 5 6 7 8 9

Meters Meters Meters Meters Meters Meters Meters Meters Meters Meters

0 −−−− 0.305 0.610 0.914 1.219 1.524 1.829 2.134 2.438 2.74310 3.048 3.353 3.658 3.962 4.267 4.572 4.877 5.182 5.486 5.79120 6.096 6.401 6.706 7.010 7.315 7.620 7.925 8.229 8.534 8.83930 9.144 9.449 9.753 10.058 10.363 10.668 10.972 11.277 11.582 11.88740 12.192 12.496 12.801 13.106 13.411 13.716 14.020 14.325 14.630 14.93550 15.240 15.544 15.849 16.154 16.459 16.763 17.068 17.373 17.678 17.98360 18.287 18.592 18.897 19.202 19.507 19.811 20.116 20.421 20.726 21.03170 21.335 21.640 21.945 22.250 22.555 22.859 23.164 23.469 23.774 24.07080 24.383 24.688 24.993 25.298 25.602 25.907 26.212 26.517 26.822 27.12690 27.431 27.736 28.041 28.346 28.651 28.955 29.260 29.565 29.870 30.174

100 30.479 30.784 31.089 31.394 31.698 32.003 32.308 32.613 32.918 33.222(TABLE I.D. 910626)

Table 3-5. Pounds to kilograms conversion

POUNDS TO KILOGRAMSCONVERSION TABLE

Pounds

0 1 2 3 4 5 6 7 8 9

Kilo-grams

Kilo-grams

Kilo-grams

Kilo-grams

Kilo-grams

Kilo-grams

Kilo-grams

Kilo-grams

Kilo-grams

Kilo-grams

0 −−−− 0.454 0.907 1.361 1.814 2.268 2.722 3.175 3.629 4.08210 4.536 4.990 5.443 5.897 6.350 6.804 7.257 7.711 8.165 8.61820 9.072 9.525 9.979 10.433 10.886 11.340 11.793 12.247 12.701 13.15430 13.608 14.061 14.515 14.969 15.422 15.876 16.329 16.783 17.237 17.69040 18.144 18.597 19.051 19.504 19.958 20.412 20.865 21.319 21.722 22.22650 22.680 23.133 23.587 24.040 24.494 24.948 25.401 25.855 26.308 26.76260 27.216 27.669 28.123 28.576 29.030 29.484 29.937 30.391 30.844 31.29870 31.751 32.205 32.659 33.112 33.566 34.019 34.473 34.927 35.380 35.83480 36.287 36.741 37.195 37.648 38.102 38.555 39.009 39.463 39.916 40.37090 40.823 41.277 41.730 42.184 42.638 43.091 43.545 43.998 44.453 44.906

100 45.359 45.813 46.266 46.720 47.174 47.627 48.081 48.534 48.988 49.442(TABLE I.D. 910648)



Table 3-6. Velocity conversion

VELOCITY CONVERSION TABLE

KNOTS MPH km/HR METERS/SEC KNOTS MPH km/HR METERS/SEC

5 5.8 9.3 2.6 105 120.8 194.4 54.010 11.5 18.5 5.1 110 126.6 203.7 56.615 17.3 27.8 7.7 115 132.3 213.0 59.220 23.0 37.0 10.3 120 138.1 222.2 61.725 28.8 46.3 12.9 125 143.8 231.5 64.330 34.5 55.6 15.4 130 149.6 240.7 66.935 40.3 64.8 18.0 135 155.4 250.0 69.440 46.0 74.1 20.6 140 161.1 259.3 72.045 51.8 83.3 23.1 145 166.9 268.5 74.650 57.5 92.6 25.7 150 172.6 277.8 77.255 63.3 101.9 28.3 155 178.4 287.0 79.760 69.0 111.1 30.9 160 184.1 296.3 82.365 74.8 120.4 33.4 165 189.9 305.6 84.970 80.6 129.6 36.0 170 195.6 314.8 87.475 86.3 138.9 38.6 175 201.4 324.1 90.080 92.1 148.1 41.2 180 207.1 333.3 92.685 97.8 157.4 43.7 185 212.9 342.6 95.290 103.6 166.7 46.3 190 218.6 351.9 97.895 109.3 175.9 48.9 195 224.4 361.1 100.3

100 115.1 185.2 51.4 200 230.2 370.4 102.9(TABLE I.D. 910629)



Table 3-7. Standard atmosphere

STANDARD ATMOSPHERE TABLE

STANDARD CONDITIONS:TEMPERATURE 15°C (59°F)PRESSURE 29.921 IN. Hg (2116.216 LB/SQ FT)DENSITY 0.0023769 SLUGS/CU FTSPEED OF SOUND 1116.89 FT/SEC (661.7 KNOTS)

CONVERSION FACTORS:1 IN. Hg = 70.727 LB/SQ FT1 IN. Hg = 0.49116 LB/SQ IN.1 KNOT = 1.151 M.P.H.1 KNOT = 1.688 FT/SEC

ALTITUDEFEET

DENSITYRATIO

σσσσ

1

√σ√σ√σ√σ

TEMPERATURE SPEED OFSOUNDKNOTS

PRESSUREIN. Hg

PRESSURERATIO°°°°C °°°°F

0 1.0000 1.0000 15.000 59.000 661.7 29.921 1.00001000 0.9711 1.0148 13.019 55.434 659.5 28.856 0.96442000 0.9428 1.0299 11.038 51.868 657.2 27.821 0.92983000 0.9151 1.0454 9.056 48.302 654.9 26.817 0.89624000 0.8881 1.0611 7.076 44.735 652.6 25.842 0.86375000 0.8617 1.0773 5.094 41.169 650.3 24.896 0.8320

6000 0.8359 1.0938 3.113 37.603 648.7 23.978 0.80147000 0.8106 1.1107 1.132 34.037 645.6 23.088 0.77168000 0.7860 1.1279 -0.850 30.471 643.3 22.225 0.74289000 0.7620 1.1456 -2.831 26.905 640.9 21.388 0.7148

10,000 0.7385 1.1637 -4.812 23.338 638.6 20.577 0.6877

11,000 0.7155 1.1822 -6.793 19.772 636.2 19.791 0.661412,000 0.6932 1.2011 -8.774 16.206 633.9 19.029 0.636013,000 0.6713 1.2205 -10.756 12.640 631.5 18.292 0.611314,000 0.6500 1.2403 -12.737 9.074 629.0 17.577 0.587515,000 0.6292 1.2606 -14.718 5.508 626.6 16.886 0.5643

16,000 0.6090 1.2815 -16.699 1.941 624.2 16.216 0.542017,000 0.5892 1.3028 -18.680 -1.625 621.8 15.569 0.520318,000 0.5699 1.3246 -20.662 -5.191 619.4 14.942 0.499419,000 0.5511 1.3470 -22.643 -8.757 617.0 14.336 0.479120,000 0.5328 1.3700 -24.624 -12.323 614.6 13.750 0.4595

21,000 0.5150 1.3935 -26.605 -15.899 612.1 13.184 0.440622,000 0.4976 1.4176 -28.587 -19.456 609.6 12.636 0.422323,000 0.4806 1.4424 -30.568 -23.022 607.1 12.107 0.404624,000 0.4642 1.4678 -32.549 -26.588 604.6 11.597 0.387425,000 0.4481 1.4938 -34.530 -30.154 602.1 11.103 0.3711

(TABLE I.D. 910646)


Table 3-8. Barometric pressure conversion

INCHES TO MILLIBARS

MERCURYINCHES

0.00 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09

MILLIBARS

28.0−−−− 948.2 948.5 948.9 949.2 949.5 949.9 950.2 950.6 950.9 951.228.1−−−− 951.6 951.9 952.3 952.6 952.9 953.3 953.6 953.9 954.3 954.628.2−−−− 955.0 955.3 955.6 956.0 956.3 956.7 957.0 957.3 957.7 958.028.3−−−− 958.3 958.7 959.0 959.4 959.7 960.0 960.4 960.7 961.1 961.428.4−−−− 961.7 962.1 962.4 962.8 963.1 963.4 963.8 964.1 964.4 964.8

28.5−−−− 965.1 965.5 965.8 966.1 966.5 966.8 967.2 967.5 967.8 968.228.6−−−− 968.5 968.8 969.2 969.5 969.9 970.2 970.5 970.9 971.2 971.628.7−−−− 971.9 972.2 972.6 972.9 973.2 973.6 973.9 974.3 974.6 974.928.8−−−− 975.3 975.6 976.0 976.3 976.6 977.0 977.3 977.7 978.0 978.328.9−−−− 978.7 979.0 979.3 979.7 980.0 980.4 980.7 981.0 981.4 981.7

29.0−−−− 982.1 982.4 982.7 983.1 983.4 983.7 984.1 984.4 984.8 985.129.1−−−− 985.4 985.8 986.1 986.5 987.8 987.1 987.5 987.8 988.2 988.529.2−−−− 988.8 989.2 989.5 989.8 990.2 990.5 990.9 991.2 991.5 991.929.3−−−− 992.2 992.6 992.9 993.2 993.6 993.9 994.2 994.6 994.9 995.329.4−−−− 995.6 995.9 996.3 996.6 997.0 997.3 997.6 998.0 998.3 998.6

29.5−−−− 999.0 999.3 999.7 1000.0 1000.4 1000.7 1001.0 1001.4 1001.7 1002.029.6−−−− 1002.4 1002.7 1003.1 1003.4 1003.7 1004.1 1004.4 1004.7 1005.1 1005.429.7−−−− 1005.8 1006.1 1006.4 1006.8 1007.1 1007.5 1007.8 1008.1 1008.5 1008.829.8−−−− 1009.1 1009.5 1009.8 1010.2 1010.5 1010.8 1011.2 1011.5 1011.9 1012.229.9−−−− 1012.5 1012.9 1013.2 1013.5 1013.9 1014.2 1014.6 1014.9 1015.2 1015.6

30.0−−−− 1015.9 1016.3 1016.6 1016.9 1017.3 1017.6 1018.0 1018.3 1018.6 1019.030.1−−−− 1019.3 1019.6 1020.0 1020.3 1020.7 1021.0 1021.3 1021.7 1022.0 1022.430.2−−−− 1022.7 1023.0 1023.4 1023.7 1024.0 1024.4 1024.7 1025.1 1025.4 1025.730.3−−−− 1026.1 1026.4 1026.7 1027.1 1027.4 1027.8 1028.1 1028.4 1028.8 1029.130.4−−−− 1029.5 1029.8 1030.1 1030.5 1030.8 1031.2 1031.5 1031.8 1032.2 1032.5

30.5−−−− 1032.9 1033.2 1033.5 1033.9 1034.2 1034.5 1034.9 1035.2 1035.5 1035.930.6−−−− 1036.2 1036.6 1036.9 1037.3 1037.6 1037.9 1038.3 1038.6 1038.9 1039.330.7−−−− 1039.6 1040.0 1040.3 1040.6 1041.0 1041.3 1041.7 1042.0 1042.3 1042.730.8−−−− 1043.0 1043.3 1043.7 1044.0 1044.4 1044.7 1045.0 1045.4 1045.7 1061.130.9−−−− 1046.4 1046.7 1047.1 1047.4 1047.8 1048.1 1048.4 1048.8 1049.1 1049.5

MILLIBARS TO INCHES

MILLIBARS

0 1 2 3 4 5 6 7 8 9

INCHES

940 27.76 27.79 27.82 27.85 27.88 27.91 27.94 27.96 27.99 28.02950 28.05 28.08 28.11 28.14 28.17 28.20 28.23 28.26 28.29 28.32960 28.35 28.38 28.41 28.44 28.47 28.50 28.53 28.56 28.58 28.61970 28.64 28.67 28.70 28.73 28.76 28.79 28.82 28.85 28.88 28.91980 28.94 28.97 29.00 29.03 29.06 29.09 29.12 29.15 29.18 29.21990 29.23 29.26 29.29 29.32 29.35 29.38 29.41 29.44 29.47 29.50

1000 29.53 29.56 29.59 29.62 29.65 29.68 29.71 29.74 29.77 29.801010 29.83 29.85 29.88 29.91 29.94 29.97 30.00 30.03 30.06 30.091020 30.12 30.15 30.18 30.21 30.24 30.27 30.30 30.33 30.36 30.391030 30.42 30.45 30.47 30.50 30.53 30.56 30.59 30.62 30.65 30.681040 30.71 30.74 30.77 30.80 30.83 30.86 30.89 30.92 30.95 30.981050 31.01 31.04 31.07 31.09 31.12 31.15 31.18 31.21 31.24 31.27

(TABLE I.D. 910647)

09 Dec 2002 Reissue 3-9/3-10


Section 4EXPANDED PERFORMANCE


Subject

TABLE OF CONTENTS

FUEL FLOW CHARTS................................................................ 4-1.................. 4-3


Page

LIST OF FIGURES

Fuel flow ..................................................................................... 4-1.................. 4-4

09 Dec 2002 Reissue 4-1/4-2



Section 4EXPANDED PERFORMANCE

4-1. FUEL FLOW CHARTS

The fuel flow charts (Figure 4-1) present fuelconsumption during level flight as a functionof altitude, OAT, airspeed, and gross weight.These charts are based on estimates andlimited flight test data. They are applicable tothe basic helicopter with all doors installedand without any optional equipment whichwould appreciably affect lift, drag, or poweravailable.

These data do not include the effects of bleedair heater, ECU, particle separator purge, or

anti-ice operation on fuel consumption. Also,fuel consumption may vary between enginesunder the same operating conditions. It isrecommended, therefore, that the operatorconduct fuel consumption checks to adjustthe presented data as necessary.

The solid line labeled LRC indicates theoptimum long range cruise airspeed for bestfuel economy. LRC is not depicted where itwo u ld o ccu r a bo ve the con t i nu o ustransmission limit line.



Figure 4-1. Fuel flow (Sheet 1 of 25)

50 60 70 80 90 100 110 120 130 140 150TRUE AIRS PEED - KNOTS

50 60 70 80 90 100 110 120 130 140INDICATED AIRSPEE D - KNOTS

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100 120 140 160 180 200 220 240 260TRUE AIRSPEED - KM /HR

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MCP TORQUE LIMIT

52.5

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M AX ENDURANCE

FUEL FLOW VS AIRSPEED

PRESSURE ALTITUDE = SEA LEVELOAT = 15°C

CLEAN CONFIGURATIONENGINE RPM 100%GENERATOR 180 AM PS

ZERO W INDHEATER OFFA NTI-ICE OFFBASIC INLET

34

M407_MD-1__FIG_4-1_(1_OF_25).EMF






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PRESSURE ALTITUDE = 2000 FTOAT = 11°C



M407_MD-1__FIG_4-1_(2_OF_25).EMF





50 60 70 80 90 100 110 120 130INDICATED AIRSPEE D - KNOTS

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MIN SPEC ENGINE + 4%

M407_MD-1__FIG_4-1_(4_OF_25).EMF






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PRESSURE ALTITUDE = 10,000 FTOAT = -5°C



VNE





M407_MD-1__FIG_4-1_(6_OF_25).EMF






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40 50 60 70 80 90 100 110INDICATED AIRSPEE D - KNOTS

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240

260

280

300

320

340

360

380

400

FU

EL

FL

OW

- L

B/H

R

25

30

35

40

45

50

55

60

65

70

75

80

85

90

TO

RQ

UE

- %


80

90

100

110

120

130

140

150

160

170

180

FU

EL

FL

OW

- K

G/H

R

GW x 1

00 LB

LRC 52.5

50

46

42

38

3430

MAX ENDURANCE









VNE

M407_MD-1__FIG_4-1_(13_OF_25).EMF






140

160

180

200

220

240

260

280

300

320

340

360

380

FU

EL

FL

OW

- L

B/H

R

25

30

35

40

45

50

55

60

65

70

75

80

85

90

TO

RQ

UE

- %


70

80

90

100

110

120

130

140

150

160

170

FU

EL

FLO

W -

KG

/HR

GW x 1

00 LB

LRC 52.5

50

46

42

38

3430

M AX ENDURANCE





VNE





M407_MD-1__FIG_4-1_(14_OF_25).EMF






140

160

180

200

220

240

260

280

300

320

340

360

380

FU

EL

FL

OW

- L

B/H

R

25

30

35

40

45

50

55

60

65

70

75

80

85

TO

RQ

UE

- %


70

80

90

100

110

120

130

140

150

160

170

FU

EL

FL

OW

- K

G/H

R

GW x 1

00 LB

LRC

52.5

50

46

42

38

3430

MAX ENDURANCE





VNE





M407_MD-1__FIG_4-1_(15_OF_25).EMF






120

140

160

180

200

220

240

260

280

300

320

340

360

FU

EL

FL

OW

- L

B/H

R

25

30

35

40

45

50

55

60

65

70

75

80

TO

RQ

UE

- %


60

70

80

90

100

110

120

130

140

150

160

FU

EL

FLO

W -

KG

/HR

GW x 100 LB

LRC

52.5

50

46

42

38

3430

MAX ENDURANCE


PRESSURE ALTITUDE = 10,000 FTOAT = 15°C



VNE





M407_MD-1__FIG_4-1_(16_OF_25).EMF






120

140

160

180

200

220

240

260

280

300

320

340

360

FU

EL

FL

OW

- L

B/H

R

25

30

35

40

45

50

55

60

65

70

75

80

TO

RQ

UE

- %


60

70

80

90

100

110

120

130

140

150

160

FU

EL

FLO

W -

KG

/HR

GW x 100 LB

LRC

52.5

50

46

42

38

34

30

MAX ENDURANCE





VNE

MIN SPEC ENGINE MCP




M407_MD-1__FIG_4-1_(17_OF_25).EMF






100

120

140

160

180

200

220

240

260

280

300

320

340

FU

EL

FL

OW

- L

B/H

R

25

30

35

40

45

50

55

60

65

70

75

TO

RQ

UE

- %


50

60

70

80

90

100

110

120

130

140

150

FU

EL

FLO

W -

KG

/HR

GW x 100 LB

LRC

52.5

50

46

42

38

34

30

MAX ENDURANCE





VNE





M407_MD-1__FIG_4-1_(18_OF_25).EMF






100

120

140

160

180

200

220

240

260

280

300

320

340

FU

EL

FL

OW

- L

B/H

R

25

30

35

40

45

50

55

60

65

70

75

TO

RQ

UE

- %


50

60

70

80

90

100

110

120

130

140

150

FU

EL

FL

OW

- K

G/H

R

GW x 100 LB

LRC

52.5

50

46

42

38

34

30

MAX ENDURANCE





VNE





M407_MD-1__FIG_4-1_(19_OF_25).EMF






160

180

200

220

240

260

280

300

320

340

360

380

400

FU

EL

FL

OW

- L

B/H

R

25

30

35

40

45

50

55

60

65

70

75

80

85

TO

RQ

UE

- %


80

90

100

110

120

130

140

150

160

170

180

FU

EL

FL

OW

- K

G/H

R

GW x 1

00 LB

LRC

52.5

50

46

42

38

30

MAX ENDURANCE





34

VNE





M407_MD-1__FIG_4-1_(20_OF_25).EMF






160

180

200

220

240

260

280

300

320

340

360

380

400

FU

EL

FL

OW

- L

B/H

R

25

30

35

40

45

50

55

60

65

70

75

80

85

TO

RQ

UE

- %


80

90

100

110

120

130

140

150

160

170

180

FU

EL

FL

OW

- K

G/H

R

GW x 100 LB

LRC

52.5

50

46

42

38

3430

MAX ENDURANCE





VNE





M407_MD-1__FIG_4-1_(21_OF_25).EMF






140

160

180

200

220

240

260

280

300

320

340

360

380

FU

EL

FL

OW

- L

B/H

R

25

30

35

40

45

50

55

60

65

70

75

80

85

TO

RQ

UE

- %

100 120 140 160 180 200 220 240 260TRUE A IRSPEED - KM /H R

70

80

90

100

110

120

130

140

150

160

170

FU

EL

FLO

W -

KG

/HR

GW x 100 LB

LRC 52.5

50

46

42

3834

30

MAX ENDURANCE



CLEA N CONFIGURATIONENGINE RPM 100%GENERATOR 180 AM PS

ZERO W INDHEATER OFFANTI-IC E OFFBASIC INLET

VNE





M407_MD-1__FIG_4-1_(22_OF_25).EMF






140

160

180

200

220

240

260

280

300

320

340

360

380

FU

EL

FL

OW

- L

B/H

R

25

30

35

40

45

50

55

60

65

70

75

80

TO

RQ

UE

- %


70

80

90

100

110

120

130

140

150

160

170

FU

EL

FL

OW

- K

G/H

R

GW x 100 LB

LRC 52.5

50

46

42

38

3430

MAX ENDURANCE





VNE





M407_MD-1__FIG_4-1_(23_OF_25).EMF






120

140

160

180

200

220

240

260

280

300

320

340

360

FU

EL

FL

OW

- L

B/H

R

25

30

35

40

45

50

55

60

65

70

75

TO

RQ

UE

- %


60

70

80

90

100

110

120

130

140

150

160

FU

EL

FLO

W -

KG

/HR

GW x 100 LB

LRC

52.5

50

46

42

38

34

30

M AX ENDURANCE





VNE





M407_MD-1__FIG_4-1_(24_OF_25).EMF






120

140

160

180

200

220

240

260

280

300

320

340

360

FU

EL

FL

OW

- L

B/H

R

25

30

35

40

45

50

55

60

65

70

75

TO

RQ

UE

- %

100 120 140 160 180 200 220 240 260TRUE A IRSPEED - KM /H R

60

70

80

90

100

110

120

130

140

150

160

FU

EL

FLO

W -

KG

/HR

GW x 100 LB

LRC

52.5

50

46

42

38

3430

M AX ENDURANCE



CLEA N CONFIGURATIONENGINE RPM 100%GENERATOR 180 AM PS

ZERO W INDHEATER OFFANTI-IC E OFFBASIC INLET

VNE





M407_MD-1__FIG_4-1_(25_OF_25).EMF

bell 407 - flight manual

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