a320_71-80v2500jarb1

ATA 71−80Engine IAE V2500

ATA 30−20Intake Ice Protection

A320 71−80V2500JARB1

EASA PART 66 B1

A 319/320/321AIRBUS

For training purposes only. Copyright by Lufthansa Technical Training.LTT is the owner of all rights to training documents and trainingsoftware.Any use outside the training measures, especially reproductionand/or copying of training documents and software − also extractsthereof −in any format all (photocopying, using electronic systemsor with the aid of other methods) is prohibited.Passing on training material and training software to third partiesfor the purpose of reproduction and/or copying is prohibited withoutthe express written consent of LTT.Copyright endorsements, trademarks or brands may not be re-moved.A tape or video recording of training courses or similar services isonly permissible with the written consent of LTT.In other respects, legal requirements, especially under copyrightand criminal law, apply.

Lufthansa Technical TrainingDept HAM USLufthansa Base HamburgWeg beim Jäger 19322335 HamburgGermany

Tel: +49 (0)40 5070 2520Fax: +49 (0)40 5070 4746E-Mail: [email protected]

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POWER PLANTINTRODUCTION

A319/A320/A321IAE V2530-A5

71-00

Page: 1FRA US/T bu Aug 2001

ATA 71 POWER PLANT

ATA 71-00 INTRODUCTIONIt is produced by International Aero Engines ( IAE ) corporation.This corporation consits of the following companys:

JAEC ( Japanese Aero Engines Corporation )Rolls RoycePratt & WhittneyMTU ( Motoren & Turbinen Union )Fiat Avio

JAEC RR P&W MTU FIAT

I A E( INTERNATIONAL AERO ENGINES )

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A319/A320/A321IAE V2530-A5

71-00

Page: 2FRA US/T bu August 2001

ENGINE MARK NUMBERSFor easy identification of the present and all future variants of the V2500, International Aero Engines has introduced a new engine designation system. All engines will retain V2500 as their generic name. The first three characters of the full designation are V25, identifying each

engine as a V2500.The next two figures indicate the engine’s rated sea − level takeoff thrust. The following letter shows the aircraft manufacturer. The last figure represents the mechanical standard of the engine.This system will provide a clear designation of a particular engine as well as asimple way of grouping by name, engines with similar characteristics.The designation V2500 − D collectively describes, irrespective of thrust, allengines for McDonnell Douglas applications and V2500 − A all engines forAirbus Industrie.Similarly, V2500 − 5 describes all engines built to the −5 mechanical standard,irrespective of airframe application.For example :The V2500 - A1 engine is used on A320 and has only a 3 stage booster.

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A319/A320/A321IAE V2530-A5

71-00


ENGINE MARK NUMBERS

V2530-A5Generic to allV2500 engines

Takeoff thrust in thousands ofpounds

Mechanical Standartsof engine

Airframe manufacturer−A for Airbus Industrie-D for McDonnellDouglas

MARK NUMBER TAKEOFF THRUST (LB) AIRCRAFT

V2500 - A1 25.000 A320 - 200

V2530 - A5 30.000 A321 - 100

V2525 - A5 25.000 A320 - 200

V2527 - A5 26.500 A320 - 200

V2528 - D5 28.000 MD - 90 - 40

V2525 - D5 25.000 MD - 90 - 30

V2522 - D5 22.000 MD - 90 - 10

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A319/A320/A321IAE V2530-A5

71-00


INTRODUCTION

The V2530 - A5 engine is a two spool, axial flow, high bypass ratio turbofanengine.

80% of the thrust is produced by the fan.

20% of thrust is produced by the engine core.Its compression system features a single stage fan, a four stage booster, and aten stage high pressure compressor. The LP compressor is driven by a fivestage low pressure turbine and the HP compressor by a two stage HP turbine.The HP turbine also drives a gearbox which, in turn, drives the engine and air-craft mounted accessories.The two shafts are supported by five main bearings.The V2500 incorporates a Full Authority Digital Electronic engine Control ( FADEC ). The control system governs all engine functions, including powermanagement. Reverse thrust is obtained by deflecting the fan airstream via ahydraulic operated thrust reverser.

IAE V2530-A5 DATA

Fan tip diameter : 63.5 in ( 161 cm )Bare engine length : 126 in ( 320 cm )Weight : 4942 lbs ( 2242 KG )Take - off thrust : 30,000 lb, flat rated to +30 deg. CBypass ratio : 5.44 : 1Overall Pressure Ratio : 31.9 :1Mass Flow lbs/s : 856 lbsN1 : 100% ( 5650 RPM )N2 : 100% ( 14950 RPM )EGT ( Takeoff ) 650 deg. CEGT ( Starting ) 635 deg. CEGT ( Max Continous/Climb ) 610 deg.C

The IAE V2530-A5 engine is flat rated.The rated thrust can be obtained for a limited time up to an ambient tem-perature of 30�C otherwise engine operating limits can be exceeded.To have a constant thrust at variable ambient conditions the engine RPMhas to be adjusted ( regulated ) to compensate the variying air density.The Thrust parameter is EPR.In case this parameter is not available theN1 is used as the Thrust parameter.

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A319/A320/A321IAE V2530-A5

71-00

Page: 5FRA US/T bu August 2001 Page: 5Figure 1 V2500 Propulsion Unit

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A319/A320/A321IAE V2530-A5

71-00


ENGINE DESCRIPTION

Gas PathA simplified view of the engine is shown below.All the air entering the engine passes trough the inlet cowl to the fan.At the fan exit the air stream divides into two flows : the core engine flow the by-pass flow

Core Engine FlowThe core engine flow passes trough the fixed inlet guide vanes to the L.P.Compressor which consits of 4 stages on the V2500 - A5 engine,then to theH.P. Compressor,the combustion section and the H.P. and L.P. turbines andfinally exhausts into the Combined Nozzle Assembly ( C.N.A. )

By-pass FlowThe fan exhaust air ( cold stream ) entering the by-pass duct passes throughthe fan outlet guide vanes and flows along the by-pass duct to exhaust into theC.N.A..

NacelleThe nacelle ensures airflow around the engine during its operation and alsoprovides protection for the engine and accessories.

The major components which comprise the nacelle are : the air inlet cowl the fan cowls ( left and right hand ) The ” C ” ducts which incorporate the hydraulically operated thrust reverser

unit. the Combined Nozzle Assembly ( CNA )

Combined Nozzle Assembly ( CNA )The core engine ” hot ” exhaust and the ” cool ” by-pass flow are mixed inthe C.N.A. before passing through the single propelling nozzle to atmosphere.

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A319/A320/A321IAE V2530-A5

71-00


V2500-A1

V2500-A5

V2500-A1

V2500-A5

BUFFER AIR COOLER OUTLET

Page: 7Figure 2 Propulsion Unit Outline

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ENGINEHAZARD AREAS

A319/A320/A321IAE V2530−A5

71-00

Page: 8FRA US/T Bu .August 2001

ATA 71-00 ENGINE HAZARD AREAS

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ENGINEHAZARD AREAS

A319/A320/A321IAE V2530−A5

71-00

Page: 9FRA US/T Bu .August 2001

ENTRY CORRIDOREXHAUST WAKE DANGERAREA 65 MPH (105 Km/h)OR LESS

INLET SUCTIONDANGER AREA

EXHAUST WAKE DANGERAREA 65 MPH (105 Km/h)OR GREATER

Page: 9Figure 3 Engine Hazard Areas

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ENGINE FUEL AND CONTROLFADEC GENERAL

A319/A320/A321IAE V2530 A5

73−20


ATA 73 ENGINE FUEL AND CONTROL

73−20 FADEC PRESENTATIONFADEC GeneralEach powerplant has a FADEC (Full Authority Digital Engine Control) system.FADEC, also called the Electronic Engine Control (EEC), is a digital controlsystem that performs complete engine management.FADEC has two−channel redundancy, with one channel active and one instandby. If one channel fails, the other automatically takes control.The system has a magnetic altemator for an internal power source. FADEC ismounted an the fan case.The Engine Interface Unit (EIU) transmits to FADEC the data it uses for enginemanagement.

FUNCTIONSThe FADEC system performs the following functions--Control of gas generator--control of fuel flow--acceleration and deceleration schedules--variable bleed valve and variable stator vane schedules control of turbine -clearance–idle settingProtection against engine exceeding limits protection against N1 and N2 over-speed monitoring of EGT during engine startPower Managementautomatic control of engine thrust rating computation of thrust parameter limitsManual management of power as a function of thrust lever position automaticManagement of power (A/THR demand).Automatic engine starting sequence control of the start valve (ON/OFF) the HP fuel valve the fuel flow the ignition (ON/OFF) monitoring of N1, N2, FF and EGT initiation of abort

and recycle (on the ground only)

Manual engine starting sequencepassive monitoring of engine control of− the start valve− the HP fuel valve− the ignitionThrust reverser control actuation of the blocker doorsengine setting during reverser operationFuel recirculation controlrecirculation of fuel to the fuel tanks according to the engine oil temperature,the fuel system configuration and the flight phase.Transmission of engine parameters and engine monitoring information to cock-pit indicatorsthe primary engine parameters the starting system statusthe thrust reverser system status the FADEC system statusDetection, isolation, and recording of failures FADEC coolingNOTE :There are no adjustments possible on the FADEC system ( e.g. Idle, Part Power etc. )

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A319/A320/A321IAE V2530 A5

73−20

Page: 11FRA US/T bu August 2001 Page: 11Figure 4 FADEC Presentation IAE V2500

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A319/A320/A321IAE V2530 A5

73−20


FADEC FUNCTIONS

The FADEC system operates compatibly with applicable aircraft systems to perform the following functions :

1 GAS generator control for steady state and transient engine operation within safe limits.

− Fuel flow control− Acceleration and deceleration schedules− Variable Stator Vane ( VSV ) and Booster Stage Bleed Valve ( BSBV ) schedules− Turbine clearance control ( HP / LP )− Idle setting.

2 Engine limits protection− Engine overspeed protection in terms of fan speed and core speed to prevent engine running over certified red lines− Engine turbine outlet gas temperature monitoring. ( EGT )

3 Power management− Automatic engine thrust rating control− Thrust parameter limit computation− manual power management through constant ratings versus throttle lever relationship

. take−off / go−around at full forward throttle lever position

. flex take−off at constant intermediate position whatever the derating is

. other ratings ( max continuous, max climb, idle, max reverse ) at associated throttle lever detent points.

− Automatic power management through direct engine power adjustment to the autothrust system demand.

4 Automatic engine start sequencing− Control of starter air valve ON / OFF− Control of HP fuel valve ( ON / OFF on ground, ON in flight )− Control of fuel schedule

− Control of ignition ( ON / OFF )− EPR, N1, N2, WF, EGT monitoring− Abort / Recycle capability on ground.

5 Thrust reverser control− Control of thrust reverser actuation ( deploying and stowing )− Control of engine power during reverser operation.- Engine idle setting during reverser transient− Control of maximum reverse power at full rearward throttle lever position.− Restow command in case of non commanded deployment.− Redeploy command in case of non commanded stowage.

6 Engine parameters transmission for cockpit indication− Primary engine parameters− Starting system status− Thrust reverser system status− FADEC system status.

7 Engine condition monitoring parameters transmission.

8 Detection, isolation, accommodation and memorization of its internalsystem failures.

9 Heat Management system (Fuel return & diverter valve control)FADEC controls the ON / OFF return to the aircraft tank in relationship with :

− Engine oil, IDG oil and fuel temperatures− Aircraft fuel system configuration− Flight phases.

Fuel Metering UnitThe fuel metering unit ( FMU ) provides fuel flow control for all operating conditions.Variable fuel metering is provided by the FMU through EEC com-mands by atorque motor controlled servo drive. Position resolvers providefeedback to the EEC. The FMU has provision to route excess fuel above en-gine requirements to the fuel diverter valve through the bypass loop.

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A319/A320/A321IAE V2530 A5

73−20


P4.9

P12,5P2,5

T2,5

FOR ENGINE TREND MONITORING

POWER

SOLENOID CONTROL VALVES

IDG

HEATERP2/T2

F FLOW

FMV FEED-BACK

� ��

THRUSTLEVER

Ignition BoxesA B

IGN BIGN A

FUEL DIVERTER & RETURN VALVE

ReturnFuel toAircraftTank

T/R REVERSER Stow / Deploy Feedback

Thrust Reverser

HydraulicPress

IGNITORS

TRUST CONTROLUNIT

RESOLVER

T/R REVERSER Stow / Deploy Command

EEC

FUEL METERINGUNIT (FMU)

HCU

ANALOG &DISCRETESIGNALS

( CH: A & B ) FEEDBACK

FUEL PRESS & COMMAND SIGNAL

FUEL FLOW TOBURNERS

HD

L B

LE

ED

VLV

‘s

COMMAND

COMMAND BY HEAT MANAGEMENT SYSTEM (HMS )

FEEDBACK

10th

7th

7th

7th

IAE V2500

(EGT)

EIU

Page: 13Figure 5 FADEC Presentation IAE V2500

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A319/A320/A321IAE V2530 A5

73−20


ENGINE CONTROL P/B’S AND SWITCHES

Engine Mode SelectorPosition CRANK :

− selects FADEC power.− allows dry and wet motoring ( ignition is not availiable ).

Position IGNITION / START :− selects FADEC power− allows engine starting (manual and auto).

Position NORM :− FADEC power selected OFF ( Engine not running )

Engine Master LeverPosition OFF :

− closes the HP fuel valve in the FMU and the LP fuel valve and resets theEEC.

Position ON :− starts the engine in automatic mode ( when the mode selector is in

IGNITION / START ).− selects fuel and ignition on during manual start procedure.

Manual Start P/B− controls the start valve (when the mode selector is in IGNITION /

START or CRANK position ).

FADEC GND PWR P/BPosition ON :

− selects FADEC power

N1 MODE P/BPosition ON :

− switches EEC from EPR Mode to N1 Mode

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A319/A320/A321IAE V2530 A5

73−20


OVERHEAD PANEL 22VU

CENTRAL PEDESTAL 115VU

MAINTENANCE PANEL 50VU

NORM

A

BC

Page: 15Figure 6 Engine Control P / B‘s and Switches

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A319/A320/A321IAE V2530 A5

73−20


49VU

2450000HMQ0

Page: 16Figure 7 Engine Circuit Breakers

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A319/A320/A321IAE V2530 A5

73−20


121VU

122VUANTI ICE

2450000UMR0

2450000VAQ0


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ENGINE INDICATINGECAM

A319/A320/A321IAE V2530−A5

77−00


ATA 77 INDICATING

77−00 ENGINE INDICATING PRESENTATION

Indication general - Primary Engine DisplayThe primary engine parameters listed below are permanently displayed on theEngine and Warning display ( E / WD ) : Engine Pressure Ratio ( EPR ) Exhaust Gas Temperature ( EGT ) N1 ( low rotor speed ) N2 ( high rotor speed ) FF ( fuel flow )

After 5 min of the power up test the indication is displayed in amber and figuresare crossed ( XX ). Normal indication can be achieved by using the FADECGRD power switches, one for each engine at the maintenace panel or by theMODE selector switch on on the Engine panel at the pedestal in CRANK orIGN / START position for both engine.If a failure occurs on any indication displayed, the indication is replaced by am-ber crosses, the analog indicator and the marks on the circle disappear, thecircle becomes amber.Only in case of certain system faults and flight phases a warning message ap-pears on the Engine Warning Display.

Secondary Engine DisplayThe lower display shows the secondary engine parameters listed below. Theengine page is available for display by command, manually or automaticallyduring engine start or in case of system fault : Total FUEL USED

For further info see ATA 73 OIL quantity

For further info see ATA 79 OIL pressure

For further info see ATA 79

OIL temperatureFor further info see ATA 79

Starter valve positions, the starter duct pressure and during eng start up,the operating Ignition system ( ONLY ON ENGINE START PAGE )

In case of high nacelle temperature a indication is provided below the en-gine oil temp. indication.

Engine Vibration − of N1 and N2 As warnings by system problems only :

− OIL FILTER COLG− Fuel FILTER CLOG− No. 4 BRG SCAV VALVE with valve position

Some engine parameters also displayed on the CRUISE page

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ENGINE INDICATINGECAM

A319/A320/A321IAE V2530−A5

77−00


FF KG / H

FOB: 19.125

� � � � � � � � � ��

PSI 35 35 PSI

IGNA BONLY ON ENGINE

START PAGE

nacc320 320

NAC temp. indication :

Page: 19Figure 9 Engine ECAM Indications

ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00

Page: 20FRA US/T-5 APR 2006

STAGE NUMBERING V2530-A5

STAGES : COMPONENT : STAGE NUMBER : NOTES :

1 FAN 1 ACOC,ACC,ACAC

1234

LOW PRESSURE COMPRESSOR

( BOOSTER )

1,52 2,32.5 B.S.B.V.

123456789

10

HIGH PRESSURE COMPRESSOR

3456789

101112

VSV ( & IGV )VSVVSVVSV

CUST. BLEED, A / I, Hdlg. Bleed,Internal Cooling

CUST. BLEED Hdlg. Bleed,

Buffer Air, 1. HPT & NGV, Muscl Air

COMBUSTION CHAMBER 20 Fuel Nozzles, 2 Ignitor Plugs

12

HIGH PRESSURE TURBINE

12

ACTIVE CLEARANCE CONTROL

12345

LOW PRESSURE TURBINE

34567

ACTIVE CLEARANCE CONTROL

COMMON NOZZLE

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ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00

Page: 21FRA US/T-5 APR 2006

V2500-A1

V2500-A5

Page: 21Figure 10 Stage Numbering

ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00


ENGINE STATIONS V2500

AERODYNAMIC STATION : STATION LOCATION : STATION USED FOR:

0 AMBIENT P0 ( ambient )

1 INTAKE / ENGINE INLET INTERFACE

2 FAN INLET Press P2 for EPR & Temp T2

12.5 FAN EXIT Press for Monitoring 12.5

2.5 L.P. COMPRESSOR ( BOOSTER EXIT ) Temp T2.5 or (CIT) & Press P2.5 for Monitoring

3 H.P. COMPRESSOR Temp T3 ( CDT ) & Press CDP ( P3 ) or BurnerPress ( Pb )

4 COMBUSTION SECTION EXIT

4.5 H.P. TURBINE EXIT

4.9 L.P. TURBINE EXIT Temp T4.9 for EGT & Press P4.9 for EPR also called P 5

5 EXHAUST

Flowpath aerodynamic stations have been established to facilitate engine per-formance assessment and monitoring.The manufacture uses numerical station designations.The station numbers areused as subscripts when designating different temperatures and pres-sures,throughout the engine.

ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00

Page: 23FRA US/T bu August 2001Figure 11 Engine Stations

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ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00


ATA 72 ENGINE

72-00 ENGINE PRESENTATION

Engine Main BearingsThe 5 bearings are located in 3 bearing compartments.

Front bearing compartmentThe front bearing compartment is located at the centre of the intermediatecase, and houses bearing No. 1, 2 & 3.

Center bearing compartment (No.4 Bearing Compartement )The center bearing compartment is located in the diffuser/combustor case andhouses bearing No. 4

Rear bearing compartmentThe rear bearing compartment is located in the turbine exhaust case No.5

The Low Pressure or N1 rotor, is supported by three bearings : Bearing 1 ( Single track thrust ball bearing ). Bearing 2 ( Single track roller bearing utilising ”squeeze film” oil damping ). Bearing 5 ( Single track roller bearing utilising ”squeeze film” oil damping ).

The High Pressure or N2 rotor is supported by two bearings : Bearing 3 ( thrust ball bearing mounted in an hydraulic damper which is

centered by a series of rod springs ( ” Squirrel Cage ” ) ). Bearing 4 ( Single track roller bearing ).

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ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00


FRONT BEAR. COMP. CENTER BEAR. COMP. REAR BEAR. COMP.

N1 BEARING NO.:

N2 BEARING NO.:

1 2 5

43

Page: 25Figure 12 Engine Bearings & Compartments

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ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00


FRONT BEARING COMPARTMENT

The bearings No. 1, 2 and 3 are located in the front bearing compartmentwhich is at the center of the intermediate module 32.The compartment is sealed using air supported carbon seals,and oil filled ( hydraulic ) seal between the two shafts. This seal is supported by 8th stageair.Adequate pressure drops across the seals to ensure satisfactory sealing .Thisis achieved by venting the compartment, by an external tube, to the de-oiler.

Gearbox DriveThe HP stubshaft, which is located axially by No 3 bearing, has at its front enda bevel drive gear which provides the drive for the main accessory gearbox,through the tower shaft.The HP stubshaft seperates from the HP compressor module at the curvic cou-pling and remains as part of the intermediate case module.

DescriptionThe drawing below shows details of No 2 and No 3 bearings.A phonic wheel is fitted to the LP stubshaft, this interacts with speed probes to provide LP shaft speed signals ( N1 ) to the EEC and theEngine Vibration Monitoring Unit ( EVMU ) which is aircraft mounted.The hydraulic seal prevents oil leakage from the compartment passingrearwards between the HP and LP shafts.No 3 bearing is hydraulically damped. The oil flow to the No. 3 bearing damperis maintained at the full oil feed pressure whilst the rest of the flow passesthrough a restrictor to drop the pressure. This allows larger jet diameters to facilitate flow tolerance control.The outer race is supported by a series of eighteen spring rods which allowsome slight radial movement of the bearing.The bearing is centralised by the rods and any radial movement is dampenedby oil pressure fed to an annulus around the bearing outer race.The gearbox drive gear is splined onto the HP shaft and retained by No 3 bearing nut.

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ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00


SealingAir

GEAR BOX DRIVE

BOOSTER AIR

PHONIC WHEEL FOR N1 RPM

SPRING ROD

Page: 27Figure 13 Front Bearing Compartment

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ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00


NO 4 BEARING COMPARTMENT

The No 4 bearing compartment is situated in a high temperature, high pressureenvironment at the centre of the combustion section.The bearing compartment is shielded from radiated heat by a heat shield andair.The No. 4 bearing compartment is cooled by 12th stage air.

12th Stage Air ( Buffer Air )This supply of cooled 12th stage air ( called ” buffer air ” ) is admitted to thespace between the chamber and first heat shield.The 12th stage air is cooledby fan air via the buffer air cooler, located on the rear left hand side of the engine.The buffer air is exhausted from the cooling spaces close to the upstream sideof the carbon seals, creating an area of cooler air from which the seal leakageis obtained.This results in an acceptable temperature of the air leaking into the bearingcompartment.Buffer air flow rates are controlled by restrictors at the outlet from the coolingpassages.

NOTE :The bearing compartment internal pressure level is determined by the area ofthe variable scavenge valve. ( called No 4 bearing scavenge valve and de-scribed in the oil system ). This valve acts as a variable restrictor in thecompartment vent / scavenge line.

NOTE :A drain hole is provided to indicate a possible leckage at the No 4 bear-ing compartment. It is located in the exhaust at 5 o clock position ( aftlooking forward )

12th Stage Air Cooler ( BUFFER AIR )The No. 4 bearing compartment air cooler is installed on the turbine casing. The exchanger is held by its coolant air duct flanges.

BUFFER AIR COOLER ( ACAC )

FAN AIR INLET

FAN AIR OUTLET

DUCTASSEMBLY

NO.4 BEARING COMPARTMENT AIRCOOLER

Page: 28

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ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00


CARBON SEALNo4 Bearing

CARBON SEAL

Spring

HEAT SHIELD

COOLED 12TH STAGE HP COMPRESSOR AIR

Page: 29Figure 14 No.4 Bearing Compartment

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ENGINEGENERAL

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REAR BEARING COMPARTMENTThe rear bearing compartment is located at the center of the LP turbine module( module 50 ) and houses No 5 bearing which supports the LP turbine rotor.The compartment is sealed at the front end by an 8th stage air supported carbon seal.At the rear is a simple cover plate, with an 0- ring and a thermally insulatedheat shield, both secured by the same twelve bolts. Inside the LP shaft there isa small disc type plug with an 0-ring seal, secured by a spring clip.There are no air or oil flows down the LP shaft.Separate venting is not necessary for this compartment because with only onecarbon seal the airflow induced by the scavenge pump gives the required pres-sure drop across the seal.The compartment is covered by an insulating heat shield.

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Page: 31FRA US/T bu August 2001 Page: 31Figure 15 Rear Bearing Compartment

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ENGINE MODULES

The engine modules are:32 the intermediate case module,31 the fan module,40 / 41 the high pressure compressor, & diffuser/combustor module,45 the high pressure turbine,50 the LP turbine60 the accessory drive gearbox.

NOTE: THE MODULE NUMBERS REFER TO THE ATA CHAPTER REF-ERENCE FOR THAT MODULE.

Fan ModuleIt consists of a single stage, wide−chord, shroudless fan and hub.Intermediate Case ModuleIt consists of the fan containment case, fan exit guide vanes ( EGV ), intermediate case, booster, low spool stubshaft, the accessory gearbox towershaft drive assembly, high spool stubshaft and the station 2.5 bleedvalve ( BSBV ). The booster consists of inlet stators, rotor assembly, andoutlet stators. The No. 1, 2 and 3 ( front ) bearing compartment is builtinto the module and contains the support bearings for the low spool andhigh spool stubshafts.

High Pressure CompressorThe HP compressor is a ten stage, axial flow module. It is comprised ofthe drum rotor assembly, the front casing which houses the variablestator vanes and the rear casing which contains the fixed stators and forms thebleed manifolds.

Diffuser / Combustor ModuleThe combustion section consists primarily of the diffuser case, annular twopiece combustor, with 20 fuel injector and 2 ignitors. The high compressor exitguide vanes and the No. 4 bearing compartment are also part of the module.The main features of the module include a close−coupled prediffuser and combustor that provide low velocity shroud air to feed the combustor liners andto minimize performance losses.

High Pressure TurbineThe high pressure turbine is a two stage turbine and drives the HP compres-sor and the accessory gearbox. Active clearance control is used to control sealclearances and to provide structural cooling.

Low Pressure TurbineThe low pressure turbine is a five stage module. Active clearance control isused to control seal clearances and to provide structural cooling.

Accessory Drive GearboxThe accessory drive gearbox provides shaft horse power to drive engine and aircraft accessories. These include fuel, oil and hydraulic pressure pumpsand electrical power generators for the EEC and for the aircraft. Thegearbox also includes provision for a starter which is used to drive the N2shaft for engine starting.

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31 - FAN 32 - INTERMEDIATE CASE

40 - HP SYSTEM41 - DIFFUSER / COMBUSTOR45 - HP TURBINE

50 - LOW PRESSURE TURBINE

60 - ACCESSORY DRIVE GEARBOX

Page: 33Figure 16 Engine Modules

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MODULE 32 INTERMEDIATE CASE

Fan CaseThe fan case provides a titanium shroud around the fan rotorand forms the outer annulus of the cold stream duct.

LP Compressor Outlet Guide VanesAerodynamic control air flow within the cold air steam duct isachieved by 60 vanes manufactured in aluminium.The vanes consist of 20 segments, each containing 3 vanes. Both sides ofthe vanes are attached to the outer and inner platforms.The outer platform is bolted to the fan case and the inner platform is pinned tothe outer shroud ring of the LP compressor stage 2.5 stator assembly.

Booster Stage bleed valve ( BSBV )The bleed valve mechanism is supported by the intermediate structure and theouter ring of the stage 2.5 vanes.Two actuating rods which are each motivated by actuators allow a axial motionto the valve ring via 2 power arms.

BOOSTER STAGE BLEED VALVE

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Fan Outlet Inner Vane Assembly

3 ea

Page: 35Figure 17 Fan Case Section

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MODULE 31 ( FAN MODULE )

Module 31 is the complete Fan assembly and comprises : 22 wide-cord ,titanium shroudless hollow fan blades 22 annulus fillers the titanium fan disc the front and rear blade retaining rings

The blades are retained in the disc radially by the dovetail root.Axial retention is provided by the front and rear blade retaining rings.Blade removal / replacement is achieved by removing the front blade retainingring and sliding the blade along the dovetail slot in the disc.The fan inner annulus is formed by 22 annulus fillers.

Nose ConeThe glass-fibre cone smoothes the airflow into the fan.It is secured to the frontblade retaining ring by 18 bolts.The nose cone is balanced during manufacture by applying weights to its insidesurface.The nose cone is unheated.Ice protection is provided by a soft rubbercone tip.The nose cone retaining bolt flange is faired by a titanium fairing which is secured by 6 bolts.

NOTE: BE CAREFUL WHEN REMOVING THE NOSE CONERETAINING BOLTS.BALANCE WEIGHTS MAY BE FITTED TO SOME OFTHE BOLTS. THE POSITION OF THE WEIGHTS MUSTBE MARKED BEFORE REMOVAL TO ENSURETHEY ARE REFITTED IN THE SAME POSITION.

Annulus FillersThe blades do not have integral platforms to form the gas−path innerannulus boundary. This function is fulfilled by annulus fillers which are locatedbetween neighbouring pairs of blades. The material of the fillers is aluminium.Each annulus filler has a hooked trunnion at the rear and a dowel pin and a pinat the front. The rear trunnion is inserted in a hole in the rear blade retainingring.The front pins are inserted in holes in the front blade retaining ring.The fillers are radially located by the front and rear blade retainingrings. Each filler is secured to the front blade retaining ring by a bolt.In order to minimize the leakage of air between the fillers and theaerofoils, there is a rubber seal bonded to each side of each filler.

Fan DiscThe fan disk is driven through a curvic coupling which attaches it tothe LP stub shaft. The curvic coupling radially locates and drives thefan disk.During manufacture of the fan disk, it is dynamically balanced byremoval of metal from a land on the disk.

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Showing Crossection ofFan Disc

Rubber Rubber

Slot Numbering

SOFT RUBBER CONE TIP

Page: 37Figure 18 LP Compressor ( Fan )

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INLET CONE REMOVALA special tool is used to remove the Inlet Cone to prevent it from damage asshown below.

NOTE: THE INLET CONE IS MADE FROM GLASSFIBER.

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A

A

Page: 39Figure 19 Inlet Cone Removal

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FAN BLADE REMOVAL / INSTALLATION

RemovalThe Nose cone is secured to the front blade retaining ring by 18 bolts.Be careful when removing the nose cone retaining bolts.Balance weights may be fitted to some of the bolts. The position of theseweights must be marked before removal to ensure they are refitted to thesame position.

The blade retaining ring is secured to the fan disc by a ring of 36 bolts. A se-cond ( outer ) ring of bolts passes through the retaining ring and screws intoeach of the 22 annulus fillers. Both rings of bolts must be removed before at-tempting to remove the front retaining ring.After all the securing bolts ( 22 + 36 ) have been removed the retaining ringcan be removed by srewing pusher bolts into the 6 threaded holes provided forthis purpose.Balance weights, if required are located on the retaining ring.

The fan blades and annulus filler positions are not identified.For this re-ason it is important to identify the blade and annulus filler position, rela-tive to the numbered slots in the fan disc, before disassembly.

Remove the annulus fillers on either side of the blade to be removed.The annulus fillers can be removed as follows : lift the front end of the annulus filler 3 to 4 inches. twist the annulus filler through about 60 deg counter - clockwise draw the annulus filler forward to clear the blades

The blade to be removed can then be pulled forward to clear the dovetail slot inthe fan disc.

InstallationAfter the new blade and the annulus fillers are fitted, The front blade retainingring can be fitted.The front blade retaining ring can only be fitted in one position which isdetermined by tree off - set locating dowells on the fan disc.

When the retaining ring is fitted to the fan disc the lettet T, etched on theretaining ring, identifies No 1 fan blade position.

NOTE: FAN BLADE INSPECTION / REPAIR ARE DESCRIBED IN THEAMM 72-31-11 PAGE BLOCK 800.

NOTE: THE MOMENT WEIGHT OF THE FAN BLADE IS WRITTEN ONTHE THE ROOT SURFACE

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”T”

MOMENT WEIGHT

Page: 41Figure 20 Fan Blade Removal / Installation

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ATA 72-31-11 FAN BLADE REPAIR

FAN BLADE INSPECTION / REPAIRBefore any repair is carried out, reference must be made to the AMM Chapter72-31-11 Page Block 800.

Repair Damage on the Low Pressure Compressor ( LPC ) Fan Blades byLocal Material Removal

CAUTION : YOU MUST USE SILICON CARBIDE TYPE ABRASIVE WHEELS,

STONES AND PAPERS TO DRESS, BLEND AND POLISH THIS COM-PONENT.

IF THE MATERIAL SHOWS A CHANGE IN COLOR, TO DARKER THAN ALIGHT STRAW COLOR, THE COMPONENT IS TO BE REJECTED.

DO NOT USE FORCE WITH MECHANICAL CUTTERS, OR THE MA-TERIAL WILL BECOME TOO HOT.

LP COMPRESSOR FAN BLADES MUST BE REPAIRED AS SOON ASDAMAGE OR WEAR IS MONITORED, TO GET BACK LP COMPRES-SOR EFFICIENCY AND EXTEND THE ROTOR BLADE LIFE.

THE MAXIMUM NUMBER OF DRESSED BLADES FOR A GIVEN THE LPCOMPRESSOR FAN BLADES SET IS THE EQUIVALENT OF THREEBLADES DRESSED TO THE MAXIMUM LIMIT. ALL THE REMAININGBLADES MUST NOT BE DRESSED.

THE MAXIMUN NUMBER OF DRESSED BLADES MUST BE OBEYED,TO PREVENT A RISK OF ENGINE VIBRATION.

PROCEDURE

NOTE: THIS REPAIR LETS YOU SCALLOP THE LEADING EDGE, RE-MOVE DAMAGE FROM THE AIRFOILSURFACE AND IF DAMAGE IS FOUND IN ZONE AD, THEN YOUMUST BLEND PARALLEL WITHTHE LEADING EDGE, TO REMOVE ANY MATERIAL ABOVE THEREPAIRED AREA BY MATERIALREMOVAL.

A. Chemically Clean the Blades ( 1 ) Use alkali cleaner (Material No. V01−300), alkani cleaner ( Material No. V01−339 ) or alkani cleaner ( Material No. V01−422 ) and prepare

the solution ( Ref. AMM TASK 70−11−50−100−010 ).

( 2 ) Wash the repaired area with a cloth soacked in the solution.

( 3 ) Use a cloth soaked in clean cold water until the area is fully cleaned.

( 4 ) If necessary repeat steps ( 2 ) and ( 3 ).

( 5 ) Wipe the area with a clean dry cloth.

B. Do a Local Penetrant Crack Test on the Damaged Blades. ( 1 ) Use fluorescent penetrant ( Material No. V06−022 ) and do a

penetrant inspection of the damaged area ( Ref. SPM 702305 ).

C. Examine the Blade Airfoil ( 1 ) Examine the blade airfoil for crack indications. Use X10 binocular under ultra violet light.

( a ) If a blade is cracked, reject it.

( 2 ) Examine the blade for damage ( Ref. TASK 72−31−11−200−010 ).

( a ) If a blade is damaged, do step ( 4.D. ) that follows.

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Page: 43FRA US/T bu August 2001 Page: 43Figure 21 Fan Blade Repair Limits

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PROCEDURE

D. Remove Local Damage on the Leading Edge ( Ref. Fig. 804 / TASK 72−31−11−991−174 )

( 1 ) Remove damage on the leading edge by removal of minimum material. Continue to remove damage until all the damage is removed. Use

portable grinding equipment.

NOTE: IF DAMAGE IS SHOWN IN ZONE AD, YOU MUST BLEND THEDAMAGE PARALLEL WITH THE BLADE LEADING EDGE, TO RE-MOVE ANY MATERIAL ABOVE THE REPAIRED AREA.

NOTE: IF YOU BLEND IN ZONE AD, YOU CAN ONLY HAVE ONE SCAL-LOP IN ZONE AC, ZONE AA AND ZONE AB, CAN EACH HAVE ASCALLOP, INDEPENDENTLY OF THE REPAIR OF ZONES AD ANDAC.

( 2 ) Remove damage as necessary on the airfoil surface by the removal of minimum material. Continue to remove damage until all the damage is removed. The maximum depth to remove the damage must not

be more than 0.015 in. ( 0.38 mm ). The diameter of the repaired area is to be 50 times the depth.

( 3 ) Make smooth the repaired area‘s. Make sure all the damaged marksare completely removed and the surface finish is made the same as the

adjacent material. Use waterproof abrasive paper ( Material No. V05−021 ), waterproof abrasive paper ( Material No. V05−020 ) and / or waterproof abrasive paper ( Material No. V05−064 ).

( 4 ) Polish the repaired area‘s, to remove scratches and make the surface finish the same as the adjacent material. Use waterproof abrasive paper ( Material No. V05−021 ), waterproof abrasive paper ( Material

No. V05−020 ) and / or waterproof abrasive paper ( Material No. V05−064 ).

NOTE: THE LAST POLISH IS TO BE IN A RADIAL DIRECTION.

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PROCEDURE

E. Examine the LP Compressor Fan Blades ( 1 ) Visually examine and measure the dimensions of the scallop on the leading edge and the airfoil surface. Make sure the maximum depth of the repair on the airfoil surfaces is not more than 0.015 in. ( 0.38 mm ). Discard the blades, if they are not in the limits specified. Use workshop inspection equipment.

F. Do a Local Penetrant Crack Test on the Damaged Blades. ( 1 ) Use fluorescent penetrant ( Material No. V06−022 ) and do a penetrant inspection of the damaged area ( Ref. SPM 702305 ).

G. Identify the Repair ( 1 ) A log book entry is necessary when you have completed this repair. Write VRS1506 in the engine log book.

( 2 ) At the next shop visit make a mark VRS1506 adjacent to the part number. Use vibro−engraving equipment.

NOTE: BLADES REPAIRED TO THIS SCHEME, MUST BE SWAB ETCHEDAND INSPECTED AS SPECIFIEDIN THE ( REF. EM 72−31−11−300−025 ) ( VRS1026 ) AND GLASSBEAD PEENED AT THENEXT SHOP VISIT, TO THE INSTRUCTIONS SPECIFIED IN THE (REF. EM 72−31−11− 300−016 )( VRS1724 ).

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MODULE 40 HP COMPRESSORThe HP compressor has 10 stages. It utilises variable inlet guide vanes at theinlet to stage 3 and variable stator vanes at stages 3, 4 and 5The front casing, which houses stages 3 to 6, is made in two halves which bolttogether along horizontal flanges. It is bolted to the intermediate casing ( module 32 ) at the front and to theouter casing at the rear.The rear compressor casing has inner and outer casings as shown. Flangeson the inner case form annular manifolds which provide 7 and 10 stage air offtakes.

NOTE: ON THE V2500-A1 THE INLET GUIDE VANES AND STAGES 3, 4, 5& 6 ARE VARIABLE.

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V2500-A1 V2500-A5

Page: 49Figure 24 HP Compressor

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COMBUSTION SECTIONThe combustion section includes the diffuser section, the combustion inner andouter liners, and the No 4 bearing assembly.

Diffuser CasingThe diffuser section is a primary structural part of the combustion section.The diffuser section has 20 mounting pads for the installation of the fuel spraynozzles. It also has two mounting pads for the two ignitor plugs.

Combustion LinerThe combustion liner is formed by the inner and outer liners.The outer liner is located by five locating pins which pass through the diffusercasing.The inner combustion liner is attached to the turbine nozzle guide vaneassembly.The inner and outer liners are manufactured from sheet metal with 100 sepa-rate liner segments attached to the inner surface. The segments can be replaced independently.

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Page: 51FRA US/T bu August 2001 Page: 51Figure 25 Combustion Section

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HP TURBINE

10th stage Make up air valveThe two position stage 10 ON / OFF valve is bolted to the 10th stage manifoldat the top of the engine compressor case.PurposeThe make up air discharges into the area around No 4 bearing housing andsupplements the normal airflows in this area and increases the cooling flowpassing to the H.P. turbine,stage 2.All of the HPT airfoils are cooled by secondary air flow.The first stage HPT blades are cooled by the HPC discharge air which flowsthrough the fist stage HPT duct assembly.The second stage vane clusters are permanent cooled by 10th stage compres-sor air mixed with thrust balance seal vent air supplied externally. The 10thstage air is supplied through 4 tubes ( 2 tubes on each engine side )Second stage HPT cooling air is a mixture of HPC discharge air and 10th stagecompressor ( make up air ). This air moves through holes in the first stageHPT air seal and the turbine front hub into the area between the hubs. The airthen goes into the second blade root and out the cooling holes,

10th Stage ” Make−up ” Air SystemIntroductionThe make up air discharges into the area around No4 bearing housing andsupplements the normal airflows in this area and increases the cooling flowpassing to the H.P. turbine, stage 2.The cooling air used is taken from the 10th stage manifold, and is controlled bya two position pneumatically operated valve.The valve position is controlled by the E.E.C. as a function of corrected N2 andaltitude.

OperationSignals from the E.E.C. will energise / deenergise the solenoid control valve.This directs pneumatic servo supplies to position the 10th stage air valve to theopen / close position.In the open position ( solenoid deenergized ) the valve allows 10th stage air toflow through two outlet tubes down the left and right hand side of the diffusercase and then pass into the engine across the diffuser area. The air then discharges into the area around No 4 bearing housing.

NOTE :The E.E.C. will keep the air valve open at all engine operating phases ex-cept cruise. The valve incorporates 2 micro switches for transmittingvalve position to the E.E.C channel A & B.The ” fail safe ” position is valve open, solenoid de−energised.

The HPC Stage 10 make up air valve and associated hardware has been de-leted from production beginning with ESN V10950

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� � � � ��

LOW OIL PRESS.SWITCH

OIL PRESS XMTR

NO.4 BEARING PRESSXMTR

EIUEEC

MIN FLOWMAX FLOW

NO.4 BEARING SCAVENGEVALVE

OIL PRESSURE

FAN AIRBUFFER AIR COOLER

( ACAC)

TO OTHER BLEED SOLENOID VALVES

EEC

TODEOILER

PB

10THSTAGESOLENOIDVALVE

REED SW

OIL AND AIR

10 TH STAGE AIR (4X)

BEARING 4COMPARTMENT

STAGE 10 AIR

BUFFERAIR

MAKE UP AIR

MAKE UPAIR VALVE

STAGE 12COMBUSTIONCHAMBER

Page: 53Figure 26 No.4 Bearing Scavenge Valve

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10 TH. STAGE MAKE UP AIR VALVEThe two position stage 10 ON / OFF valve is bolted to the 10th stage manifoldat the top of the engine compressor case.The valve is equipped with a position indicator ( closed or open )

The HPC Stage 10 make up air valve and associated hardware has been de-leted from production beginning with ESN V10950

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POSITION INDICATORVISUAL

10th STAGE PRESSTO NO4 BEARING SCAVENGE VALVE

2 POSITION FEEDBACK SWITCESTO EEC

P3 SERVOPRESS.

AIR OUTLET TUBES

O

C

INDICATOR PIN

SOLENOID

CONTROL

VALVE

Page: 55Figure 27 Stage10 to HPT Air Control Valve

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COMMON NOZZLE ASSEMBLY (CNA) GeneralThe mixed exhaust system collects two flows of air.The first is the cold airflow, which is the fan bypass air.The second is the hot airflow which comes from the engine core.The mixed exhaust system is made up of the common nozzle exhaust collectorand the engine exhaust cone. The common exhaust collector admits the hot and cold gas outflows. These

gas outflows then go out to the atmosphere through the common nozzle. The nozzle forms a convergent duct which increases the speed of the mixed

gas to give forward thrust. The engine exhaust cone forms the inner contour of the common nozzle

exhaust collector. It is made of a welded inco 625 honeycomb perforatedpanel for sound attenuation, an attachment ring and a closure panel.

Interface seals provide sealing between the exhaust collector, the thrust reverser and the pylon .

The cold airflow exhaust is part of the thrust reverser system described in 78−30−00. When the thrust reverser operates, the cold and hot outflowsdivide, and go in different directions.

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Page: 57FRA US/T bu August 2001 Page: 57Figure 28 Common Nozzle Assemply

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ATA 72−60 ACCESSORY DRIVE GEARBOX

ANGLE AND MAIN GEARBOX

The cast aluminium gearbox assembly transmits power from the engine to pro-vide drives for the accessories mounted on the gearbox front and rear faces.During engine starting the gearbox also transmits power from the pneumaticstarter motor to the engine.The gearbox also provides a hand cranking for the HP rotor ( N2 ) formaintenance operations.The gearbox is mounted by 4 flexible links to the bottom of the fan case.main gearbox 3 linksangle gearbox 1 link

Features :

Front Face Individually replaceable drive units Magnetic chip detectors Main gearbox 2 magnetic chip detectors Angle gearbox 1 magnetic chip detector De−oiler Pneumatic starter Dedicated generator / alternator Hydraulic pump Oil Pressure pump

Rear Face Fuel pumps ( and Fuel Metering Unit FMU ) Oil scavenge pumps unit Integrated Drive Generator System ( I.D.G.)

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REAR VIEW

FRONT VIEW

Manual Drive

SCAVENGE

& PRESS FILTER

Page: 59Figure 29 Angle and Main Gearbox

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DRIVE SEAL

The sealol sealThe picture below shows a typical SEALOL SEAL ( carbon drive seal ) installation ( Starter ).This type of seals are used on the drive pads on the gearbox.Consists of the following parts : A mating ring ( glazed face ) with four lugs engageing the four correspond-

ing slots in the gearshaft ball bearing. A cover, secured to the bearing housing with nuts, to ensure constant

contact between the glazed face and the static part of the seal.The sealol seals are matched assemblies. If one of the components isdamaged, replace the complete seal !

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ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00


SEALOL SEAL

Page: 61Figure 30 Drive Seals

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ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00

Page: 62FRA US/T bu August 2001 Page: 62Figure 31 Engine Components Location (L/H side)

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ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00

Page: 63FRA US/T bu August 2001 Page: 63Figure 32 Engine Components Location (R/H side)

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ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00


ENGINE FLANGESFlanges are located on the engine for attachment of brackets,claps,bolt,etc.

Physical DescriptionThe external flanges of the engine have been assigned letter designations al-phanumerical from A to U.The letters I,O and Q are not used.The letter desig-nations are used for flange identification whenever it is necessary to be explicitabout flange location.

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ENGINEGENERAL

A319/A320/A321IAE V2530-A5

72-00

Page: 65FRA US/T bu August 2001 Page: 65Figure 33 Engine Flanges

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ENGINEBORESCOPING

A319/A320/A321IAE V2530-A5

72-00


ATA 72-00 BORESCOPING

GENERAL

Hand CrankingA access to crank the HP compressor manually is provided at the front face ofthe gearbox between the Starter and the deticated alternator.

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ENGINEBORESCOPING

A319/A320/A321IAE V2530-A5

72-00

Page: 67FRA US/T bu August 2001 Page: 67Figure 34 Manual Handcranking

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ENGINEBORESCOPING

A319/A320/A321IAE V2530-A5

72-00


BORESCOOPING GENERAL

Hand CrankingA access to crank the HP compressor manually is provided at the front face ofthe gearbox between the Starter and the deticated alternator.

BORESCOOPE INSPECTION OF THE HP COMP.Borescope ports are provided to give acess for visual inspection of the compressor and the turbine . For furter information and limits refer to AMM 72-00-00.

Inspection/Check Procedure Install the tool to turn the HP system. Prepare the borescope equipment for use as given in the makers

instructions. Carefully put the borescope probe into the access port of the stage of the

compressor you want to examine .

NOTE: USE AN 8MM PROBE FOR PORTSX,A,B AND A 5.5MM PROBEFOR PORTS C,D,E,F & G AND A FLEXIBLE BORESCOPE FORINSPECTION OF THE HEATSHIELD ASSEMBLIES.

Whilst turning the HP system, examine each blade in turn for:− Nicks & Tears− Cracks− Dents− Tip damage & discolouration

NOTE: BLADE NUMBERS & DIMENSIONS ARE SHOWN FOR EACHSTAGE.

Examples of blade damage limits are in AMM 72-00-00 On completion of the inspection remove the borescope probe from the en-

gine and refit the access port covers as described on the next page. Remove the tool used to turn the HP system & return the engine to normal.

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ENGINEBORESCOPING

A319/A320/A321IAE V2530-A5

72-00


STAGE OF COMPRESSORTO BE EXAMIND

ACCESS PORTTO BE USED

3 to 44 to 67 to 88 to 99 to 10

11 to 12

ABDEFG

NOTE: Port ”B” is available at both sides of the engineThe left hand side is better accessible

V2530-A5V2500-A1

V2530-A5

Page: 69Figure 35 HP Compressor Borescope Access

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ENGINEBORESCOPING

A319/A320/A321IAE V2530-A5

72-00


BORESCOPE INSPECTION OF THE HP COMP. CONT.

Borescope Access

NOTE: IAE RECOMMENDS THAT ONLY THE STAGE 3 & 12 HP COM-PRESSOR BLADES ARE EXAMINED WITH THE ENGINE ON−WING.

NOTE: ACCESS PORT D SHOULD NOT BE USED ON ENGINES THATARE PRE SBE72−0033 AS DAMAGE CAN BE CAUSED TO THEBORESCOPE EQUIPMENT.

Remove the required borescope access part covers X,A,B,C,D,E,F,G, byremoving the attaching bolts. The diagram below shows which stage are accessed through each port.

Remove the old jointing compound from around the access ports and ac-cess port covers using a non−metallic scraper and a lint free cloth mademoist with cleaning fluid.

Prior to installation of the borescope access port covers it Is necessary toapply jointing compound. The procedure to be taken is:

Access ports X, A, B & C− Apply a thin layer of jointing compound to the mating faces using a stiff

bristle brush. Do not apply within 0.12 to 0.16in (3 to 4mm) of accessport.

− Wait 10 minutes, install access port cover & attach with bolts. Torqueload to between 85 − 105 lbf in.

− Re−torque again to same figures after 2 minutes then remove excessjointing compound.

Access ports D,E,F & G.− Do not require jointing compound.

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ENGINEBORESCOPING

A319/A320/A321IAE V2530-A5

72-00


STAGE OF COMPRESSORTO BE EXAMIND

ACCESS PORTTO BE USED

3 to 45 to 6

BC

VIGV TO 3 -LE X

Page: 71Figure 36 HP Compressor Borescope Access

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ENGINEMOUNTS

A319/A320/321IAE V2530−A5

71-20

Page: 72FRA US/T Bu July 01

ATA 71 POWER PLANT71-20 ENGINE MOUNTS

GeneralThe engine mounts support the engine by transmitting loads from the enginecase to the pylon structure.They allow thermal expansion of the engine without inducing additional loadinto the mount system.Each engine mount design provides dual load paths to ensure safe operationif one member fail.

The engine/pylon connection is achieved by means of a two−mount system :− the forward mount : it is attached to the engine via the intermediate casing. It takes the X loads(thrust), Y loads (lateral) and Z loads (vertical).− the aft mount :it is attached to the engine via the exhaust casing. It takes the loads in a planenormal to the engine centerline i.e.: Y loads (lateral), Z loads (vertical) and Mx(engine rotational inertia moment + Y load transfer moment).

Component LocationThe front mount is installed at the top center of the low pressure compressor case.The rear mount is installed at the top center of the low pressure turbinecase.The engine mount system has these components:

− A front mount− A rear mount .

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ENGINEMOUNTS

A319/A320/321IAE V2530−A5

71-20

Page: 73FRA US/T Bu July 01 Page: 73Figure 37 Mounts and Loads

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POWER PLANTENGINE MOUNTS

A319/A320A321IAE V2530-A5

71-20


71-20 ENGINE MOUNTS

GeneralThe engine is attached to the aircraft pylon by two mount assemblies, one at the front and one at the rear of the engine.The mount assemblies transmitloads from the engine to the aircraft structure.Spherical bearings in each mount permit thermal expansion and somemovement between the engine and the pylon.Both mounts are made to be fail−safe and have a tolerance to damage.

FORWARD ENGINE MOUNTThe front mount has these parts: Two thrust links. A beam assembly. A cross beam assembly. A support bearing assembly.

The thrust links attach to lugs on the cross beam and to the engine mount lugson the low pressure compressor using solid pins. A spherical bearing isinstalled at each end of the links.Vertical and side loads are transmitted throughthe support bearing to the beam assembly and then to the aircraft pylon.The beam assembly is aligned on the aircraft pylon by two shear pins and at-tached with five bolts.The thrust of the engine is transmitted through the thrust links, the cross beamassembly and the beam assembly to the aircraft pylon.The support bearing permits the engine to turn so that torsional loads are nottransmitted to the aircraft structure.The front mount is made to be fail−safe. If one of the two thrust links or thecross beam should fail, then thrust loads are transmitted through the ball stopand into the beam assembly. The thrust is then transmitted to the pylon struc-ture.

AFT ENGINE MOUNTThe aft mount has these parts: Two side links. A center link. A beam assembly.

The two side links attach to the beam assembly at one end and the engine aftmount ring on the low pressure turbine case at the other end.The aft mount is aligned on the pylon by two shearpins and is attachedto the pylon by four bolts and washers.Vertical and side loads are transmitted through the side links andbeam assembly and into the pylon.Torsional loads are transmitted by the center link to the beamassembly and in to the pylon.The mount is made to be fail−safe. The side links are each made up of twoparts which are attached together to make one unit. If one part of the linkshould fail, the remaining part will transmit the loads to the beam assembly.

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POWER PLANTENGINE MOUNTS

A319/A320A321IAE V2530-A5

71-20


Thrust Link

Thrust Link

Support Bearing

Cross Beam Assembly

Pylon Mount

Fail Safe Bolt

Beam Assembly SHEAR PINS

AFT MOUNTFORWARD MOUNT

Page: 75Figure 38 Engine Mounts

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POWER PLANTCOWLINGS

A319/A320/A321V2530-A5

71-10

Page: Page: 76FRA US/T bu August 2001

ATA 71-10 NACELLE ACCESS DOORS & OPENINGSNACELLE GENERALThe nacelle ensures airflow around the engine during its operation and alsoprovides protection for the engine and accessories.

The major components which comprise the nacelle are: the air inlet cowl the fan cowls (left and right hand) The ”C” ducts which incorporate the hydraulically operated thrust reverser

unit. the Combined Nozzle Assembly (CNA)

ACCESS DOORS & OPENINGS

Access to units mounted on the low pressure compressor (fan) case and ex-ternal gearbox is gained by opening the hinged fan cowls.Access to the core engine ,and the units mounted on it ,is gained by openingthe hinged ”C” ducts.

Pressure relief Doors:Two access doors also operate as pressure relief doors.They are installed oneach nacelle. The air starter valve and pressure relief door in the right fan cowl and the oil tank service pressure relief door in the left fan cowl.

The two pressure relief doors protect the core compartment against a differen-tial overpressure of 0.2 bar (2.9007 psi) and more.Spring−loaded latches hold the doors in place. If overpressure causes one orthe two doors in a nacelle to open during flight, they will not latch close againautomatically. The door (doors) will be found open during ground inspections.

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POWER PLANTCOWLINGS

A319/A320/A321V2530-A5

71-10


RIGHT SIDE

LEFT SIDE

ACAC OUTLET

ACAC OUTLET

STRAKE

STRAKE

PRESSURERELIEF DOOR

Page: 77Figure 39 Nacelle Access Doors

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POWER PLANTCOWLINGS

A319/A320/A321V2530-A5

71-10


FAN COWLS OPENING / CLOSING

The fan cowl doors extend rearwards from the inlet cowl to overlap leadingedge of the ”C” ducts.When in the open position the fan cowls are supported bytwo telescopic hold − open struts,using support points provided on the fan case(rear) and inlet cowl (front). Storage brackets are provided to securely locatethe struts when they are not in use.

WarningThe fan cowl hold open struts must be in the extended position and bothstruts must always be used to hold the doors open.

Be careful when opening the doors in winds of more than 26 knots(30mph)

WarningThe fan cowl doors must not be opened in winds of more than 52 knots(60mph)

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POWER PLANTCOWLINGS

A319/A320/A321V2530-A5

71-10


DETAIL AT 4 POSITIONS

Page: 79Figure 40 Fan Cowls Opening / Closing

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Power PlantGeneral

A319/A320/A321V2530-A5

71-10


FAN COWL LATCH ADJUSTMENTThe mismatch between the two cowl doors can be adjusted by fitting / remov-ing shims,as shown below.Latch tension is adjusted by use of the adjusting nut at the back of the latchkeeper as shown below.

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Power PlantGeneral

A319/A320/A321V2530-A5

71-10

Page: Page: 81FRA US/T bu August 2001 Page: 81Figure 41 Fan Cowl Latch Adjustment

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ENGINE EXHAUSTTHRUST REVERSER COWLS

A319/A320/A321V2530-A5

78-32


ATA 78-32 THRUST REVERSER COWL DOORST/R COWLING ( ”C-DUCT” ) OPENING / CLOSING

Caution

Before opening:

1. Wing slats must be retracted and deactivated.

2. All 6 latches & take - up devices must be released.

3. If reverser is deployed, pylon fairing must be removed.

4. Deactivate Thrust Reverser Hydraulic Control Unit ( HCU )

5. FADEC power ”OFF”

6. Put Warning Notices in the Cockpit

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A319/A320/A321V2530-A5

78-32


FAIRING

With deployed reverser thefairing must be removed !

PYLON

REVERSER CASCADES

Page: 83Figure 42 C-Duct Opening/Closing

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A319/A320/A321V2530-A5

78-32


THRUST REVERSER HALF LATCHES

6 Latches are provided to keep the Thrust Reverser Halfs in the closed position.They are located : 1 Front latch ( access through the left fan cowl ) 3 Bifurcation latches ( access through a panel under the C-Duct halves ) 2 latches on the reverser translating sleeve ( Double Latch )

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A319/A320/A321V2530-A5

78-32


A B

C

Page: 85Figure 43 Thrust Reverser Half Latches

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A319/A320/A321V2530-A5

78-32


LATCH ACCESS PANEL & TAKE UP DEVICEAn access panel ,as shown below , is provided to gain access to the three BIFURCATION ”C” duct latches and the ”C” duct take up device (also called,Auxiliary Latch Assembly ).The take up device is a ”turnbuckle” arrangement which is used to draw thetwo ”C” ducts together.This is necessary to compress the ”C” duct seals farenough to enable the latch hooks to engage with the latch keepers.The take up device is used both when closing and opening the ”C”ducts.The take up device must be disengaged and returned to its stowage bracket,in-side the L/H ”C” duct,when not in use.

NOTE: RED OPEN FLAGS ,INSTALLED ON THE C-DUCT INDICATETHAT THE BIFURCATION LATCHES ARE OPEN.

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A319/A320/A321V2530-A5

78-32


DETAIL VIEW of a typicalLatch - Open Indicatoron the Bifurcation Latch.

Open-Indicator( 3 installed )

Page: 87Figure 44 Latch Panel & Take Up Device

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A319/A320/A321V2530-A5

78-32


FRONT LATCH AND OPEN INDICATOR

Access to the front latch is gained through the left hand fan cowl. The latch isequipped with a red open indicator.The open -indicator gets in view through a gap in the cowling ( also when thethrust reverser halfs are closed ) to indicate a not propper closed reverser cowl.

CAUTION: MAKE SURE THAT YOU POSITION THE FRONT LATCHCORRECTLY AGAINST THE FRONT LATCH OPEN INDICA-TOR WHILE YOU PULL THE THRUST REVERSER HALVESTOGETHER WITH THE AUXILIARY LATCH ASSEM-BLY.(TAKE UP DEVICE)IF YOU DO NOT DO THIS ,THE FRONT LATCH CAN GETCAUGHT BETWEEN THE THRUST REVERSER HALVESAND THE AUXILIARY LATCH ASSEMBLY AND THE HOOKCAN GET DAMAGED.

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A319/A320/A321V2530-A5

78-32


RED FRONT LATCHOPEN INDICATOR FRONT LATCH

FRONT LATCH OPENINDICATOR

SPRING

B

Page: 89Figure 45 Front Latch with Open Indicator

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A319/A320/A321V2530-A5

78-32


C - DUCT OPENING / CLOSING SYSTEMOn each ”C” duct a single acting hydraulic actuator is provided for opening.A hydraulic hand pump must be connected to a self sealing /quick release hy-draulic connection for opening.

NOTE: THE HYDRAULIC FLUID USED IN THE SYSTEM IS ENGINE LU-BRICATING OIL.

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A319/A320/A321V2530-A5

78-32

Page: 91FRA US/T bu August 2001 Page: 91Figure 46 ”C” Duct Opening/Closing

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A319/A320/A321V2530-A5

78-32


C - DUCT HOLD OPEN STRUTSTwo hold open struts are provided on each C - duct to support the C - ducts inthe open position.The struts engage with anchorage points located on the engine as shown be-low.When,not in use the struts are located in stowage brackets provided inside theC - ductThe front strut is a fixed length strut.The rear strut is a telescopic strut and must be extended before use.The arrangement for the L.H. ’C’ duct is shown below, the R.H. ’C’ duct issimilar.

WARNING: BOTH STRUTS MUST ALWAYS BE USED TO SUPPORTTHE ’C’ DUCTS IN THE OPEN POSITION. THE ’C’ DUCTSWEIGH APPROX 578 LBS EACH. SERIOUS INJURY TOPERSONNEL WORKING UNDER THE ’C’ DUCTS CAN OC-CUR IF THE ’C’ DUCT IS SUDDENLY RELEASED.

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A319/A320/A321V2530-A5

78-32

Page: 93FRA US/T bu August 2001 Page: 93Figure 47 „C“ Duct Hold Open Struts

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ENGINEOIL SYSTEM

A319/A320/A321IAE V2530-A5

79−00


ATA 79 OIL

79−00 GENERAL

Oil System PresentationThe lubrication system is self−contained and thus requires no airframe suppliedcomponents other than certain instrumentation and remote fill and drain portdisconnectors on the oil tank.These ports are used to refill the oil tank promptlyand precisely by allowing the airlines to quick−connect a pressurized oil lineand a drain line.

Lubrication System ComponentsThe lubrication system consits of four subsystems:− the lubrication supply system− the lubrication scavenge system− the oil seal pressurization system− the sump venting system.The oil system lubricates the engine components. It contains − the oil tank − the lube and scavenge pump modules − the fuel/oil heat and air/oil heat exchangers − the filters, chip detectors, pressure relief and bypass valves.

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ENGINEOIL SYSTEM

A319/A320/A321IAE V2530-A5

79−00

Page: Page: 95FRA US/T bu August 2001 Page: 95Figure 48 Oil System Basic Schematic

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ENGINEOIL SYSTEM

A319/A320/A321IAE V2530-A5

79−00


79−00 GENERAL

Oil System Presentation

System DescriptionThe lubrication system is self−contained and thus requires no airframe suppliedcomponents other than certain instrumentation and remote fill and drain portdisconnectors on the oil tank.These ports are used to refill the oil tank promptlyand precisely by allowing the airlines to quick−connect a pressurized oil lineand a drain line.It is a hot tank system that is not pressure regulated.Oil from the oil tank enters the one stage pressure pump and the dischargeflow is sent directly to the oil filter. A coarse cleanable filter is employed.The oil then is piped through the air cooled oil cooler and the fuel cooled oilcooler ,which are part of the Heat Management System (HMS) ,which ensuresthat engine oil,IDG oil and fuel temperatures are maintained at acceptable lev-els, to the bearings.Except for the No 3 bearing damper and the No.4 bearingcompartment,the pressure supplied to each location is controlled by a restric-tor.There is a ”last chance” strainer at the entry of each compartment to pre-vent blockage by any debris / carbon flakes in the oil.The savenge oil is then piped,either directly or through the de-oiler to the 5stage scavenge pumps.There is a disposable cartridge type scavenge filter atthe outlet of the scavenge pumps before returning to the oil tank.A valve allowsoil to bypass the scavenge filter when the filter differential pressure exceeds 20psi. A differential pressure warning switch set at 12 psi, gives cockpit indicationof impending scavenge filter bypass.The oil pressure is measured as a differential between the main supply linepressure, upstream of any restrictors, and the pressure in the No.4 bearingcompartment scavenge line, upstream of the two position scavenge valve. A low pressure warning switch, which is set for 60 psi, is provided in the mainoil line before the bearing compartments and after the ACOC and FCOC at thesame tapping points as the oil pressure sensor.This allows for cockpit monitor-ing of low oil pressure.The engine oil temperature is measured in the combinedscavenge line to the oil tank.The No.4 bearing two position scavenge valve is operated pnuematically bytenth stage air and controls vented air flow from the bearing compartment inresponse to specific levels of engine thrust setting.At engine idle power, the-valve opens to provide the maximum area for scavenge flow. At higher power,

the valve closes to a reduced area which provides,adequate pressure in theNo.4 bearing compartment to protect the seals by maintaining low pressuredifferentials across compartment walls and minimizes air leakage into the bear-ing chamber.The scavenge valve pressure transducer senses the pressure present in thescavenge line upstream of the scavenge valve and supplies a signal to the EIU.A pressure relief valve at the filter housing limits pump discharge pressure toapproximately 450 psi to protect downstream components.Lubrication System ComponentsThe lubrication system consits of four subsystems:− the lubrication supply system− the lubrication scavenge system− the oil seal pressurization system− the sump venting system.

System Monitoring and LimitationsThe operation of the engine oil system may be monitored by the following flightdeck indications. engine oil pressure engine oil temperature

− MINIMUM STARTING: - 400 C− MIN.PRIOR EXCEEDING IDLE : -100C− MIN. PRIOR TAKE OFF: 500C− MAX CONTINIOUS: 1550C− MAX TRANSIENT: 1650C

oil tank contents 25 US quartsIn addition warnings may be given for the following non normal conditions: low oil pressure

− RED LINE LIMIT: 60 PSI− AMBER LINE LIMIT: 80 PSI

scavenge filter clogged. No. 4 compartment scavenge valve inoperative.

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ENGINEOIL SYSTEM

A319/A320/A321IAE V2530-A5

79−00


REED SWITCH

ANTI-DRAINVLV

FLOW TIMING VLV

OIL QTY XMTR

FUEL FILTER

FCOC ENG OIL

ACOC

NO 1, 2 & 3 BEARINGS

SCAVENGE FILTER ∆ P SWITCH ( 12 PSI , ECAM MESS: ” OIL FILTER CLOG )”

OIL TEMPERATURESENSOR( HMS )

BYPASSVLV‘ S

SCAVENGE FILTER

OILPRESS. XMTR

LOW OILPRESS. WARNINGSWITCH( 60 PSI )

NO 4 BEARINGCOMPARTMENT 2 POSITIONSCAVENGE VLV

10TH

STAGEAIR

COLD START PRESSRELIEF VLV( 450 ∆ PSI )

FAN AIR

FUEL

BUFFERAIR ( 12TH ) CAVITY DRAIN LINE

BIFURCATION PANEL

DE-OILER

SCAVENGEFILTER BYPASS VLV( 20 PSI ∆ P )

RESTRICTOR

SCAVENGE PUMPS

OIL TANK PRESSURIZATION VLV

NO 4 BEARING PRESS XMTR

NO.5 BEARING

NO. 4 BEARING

IN

OUT

MASTER CHIP DETECTOR

OIL TEMPERATURE SENSOR

BREATERAIR

OIL TANK FILLER CAP

Page: 97Figure 49 Oil System Schematic

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OIL SYSTEMINDICATING

A319/A320/A321IAE V2530-A5

79-30

Page: 98FRA US/T Bu August 2001

79-30 OIL INDICATING SYSTEM

GeneralThe oil system monitoring is performed by:

- indications: oil quantity (quarts) oil temperature (degree celsius) oil pressure (psi)

- audio and visual warnings: oil low pressure (LO PRESS) oil filter clogging (OIL FILTER CLOG)

ECAM OIL INDICATIONS1.- Oil quantity indication flashes green (Advisory): when QTY <4quarts.

2.- Oil pressure indication color turns red (Warning) : when press <60PSI.

3.- Oiltemperature indication flashes green (Advisory) : when TEMP >156 deg.C turns amber when oil TEMP < 10 deg C or > 165 deg C.

Oil HI TEMP is displayed : when oil TEMP >165 deg C or 156 deg C more than 15 min.

4.- Oil filter clog (White & amber) warning appears on the screen when the engine scavenge filter is clogged.

5.−Eng.1 (2) BEARING 4 OIL SYS. ( class 2 ) and a message SCAVENGE VALVE FAULT is displayed when the valve is

not in the correct position according to the sensed burner pressure. The massage HI PRESS is displayed when the No. 4 bearing compartment

pressure is is to high according to the valve position and a high burnerpress.(possible Carbon seal failure ) or scavenge valve stuck in closed orscavenge line pressure sensor malfunction.

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79-30


1

2

3

Page: 99Figure 50 ECAM Oil Indication

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A319/A320/A321IAE V2530-A5

79-30


OIL QUANTITY INDICATINGThe analog signal from the oil quantity transmitter is sent to:

− the SDAC1− the SDAC2− the EIU which transforms the analog signal into a digital signal.

The DMC’s process the information received as a priority order fromthe EIU’s through FWC 1 and 2, SDAC1, SDAC2.The oil quantity displayed in green on the ECAM display unit is graduated from: 0 to 25.8 qts in analog form (the normal max-usable oil quantity in the tank

is 25 US qts,,the maximum oil tank capacity is 30.5 US qts) 0 to 99.9 in digital form.

OIL TEMPERATURE INDICATION

The analog signal from the scavenge oil temperature thermocouple is trans-mitted to the EIU.The EIU transforms this signal into a digital signal.This digital signal is then transmitted to the lower ECAM display unit throughthe FWCs and the DMC.The ECAM oil temperature indication scale is graduated from 0 deg.C to999 deg.C .

OIL PRESSURE INDICATIONThe analog signal from the oil pressure transmitter is transmitted to the SDAC1,SDAC2 and the EIU .The EIU transforms this signal into a digital signal.This digital signal is then transmitted to the lower ECAM display unit throughthe FWCs and the DMC.The order of priority has been defined as follows:

SDAC 1SDAC 2EIU.

The oil pressure indication scale is graduated from 0 - 400 PSI .

LOW OIL PRESSURE SWITCHThe low oil pressure information is send to different aircraft systems.

Low Oil Pressure switching: To Steering (ATA 32-51) To Door Warning (ATA 52-73) To FWC (ATA 31-52) To FAC (ATA 22 ) To FMGC (ATA 22-65) To IDG System Control (ATA 24-21 )

Low Oil Pressure Switching via EIU: To CIDS (ATA 23-73) To DFDRS INTCOM Monitoring (ATA 31-33 ) To CVR Power Supply (ATA 23-71) To WHC (ATA 30-42) To PHC (ATA 30-31) To FCDC (ATA 27-95) To Blue Main Hydraulic PWR (ATA 29-12) To Rain RPLNT ( ATA 30-45 )

SCAV. FILT. DIFF. PRESSURE WARNINGThe Scavenge filter diff.pressure warning is send to the SDAC 1,2 and then toECAM. A message will be displayed on the E/WD.

NO.4 BEARING WARNINGTwo EIU logics provide a warning message to the ECAM :

Eng.1 (2) BEARING 4 OIL SYS. ( class 2 )and a message SCAVENGE VALVE FAULT is displayed when the valve isnot in the correct position according to the sensed burner pressure.The massage HI PRESS is displayed when the No. 4 bearing compartmentpressure is is to high according to the valve position and a high burnerpress.(possible Carbon seal failure ) or scavenge valve stuck in closed orscavenge line pressure sensor malfunction.

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79-30

Page: 101FRA US/T Bu August 2001 Page: 101Figure 51 Basic Schematic

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ENGINEOIL SYSTEM

A319/A320/A321IAE V2530-A5

79−00


OIL TANKThe tank is located on the top L. H. side of the gearbox.The normal max-usable oil quantity in the tank is 25 US qts,,the maximum oiltank capacity is 30.5 US qtsFeatures: oil qty. transmitter pressure and gravity fill ports sight glass for level indication internal deaerator tank pressurisation valve ( 6 psi ) strainer in tank outlet mounting for scavenge filter and master chip detector

ENGINE OIL SERVICING

Where conditions permit,the oil tank should be checked and oil added,if neces-sary , within a period of 5 to 20 minutes after engine shutdown.If the engine isstopped for 10 hours or more,a dry motoring must be performed.This makesure that the oil level shown in the tank is correct before oil is added.

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SIGHT GLASS

Page: 103Figure 52 Oil Tank

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ENGINEOIL SYSTEM

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79-00 OIL SYSYSTEM COMPONENTS

Oil TankThe tank is located on the top L. H. side of the gearbox.The normal max-usable oil quantity in the tank is 25 US qts,,the maximum oiltank capacity is 30.5 US qtsFeatures: oil qty. transmitter pressure and gravity fill ports sight glass for level indication internal deaerator tank pressurisation valve ( 6 psi ) strainer in tank outlet mounting for scavenge filter and master chip detector

Engine Oil ServicingWhere conditions permit,the oil tank should be checked and oil added,if neces-sary , within a period of 5 to 20 minutes after engine shutdown.If the engine isstopped for 10 hours or more,a dry motoring must be performed.This makesure that the oil level shown in the tank is correct before oil is added.

OIL QUANTITY TRANSMITTERThe oil quantity transmitter is located in the oil tank.

Power SupplyThe system is supplied with 28VDC from busbar ENG 1,101PP (DC BUS 1 )through circuit breaker 1EN1 (2EN1).

Description :The oil quantity tranmitter is a tank probe with a capacitor (tube portion) and anelectronic module (on the top of the transmitter) for probe energizing and signaloutput.

Output voltage :1VDC to 9VDC varying linearly with the usable oil quantity from 0 to 25.8quarts.

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79−00


SIGHT GLASS

A

A

OIL QUANTITY TRANSMITTER

Page: 105Figure 53 Oil Tank

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ENGINEOIL SYSTEM

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79−00


OIL PRESSURE PUMPThe pressure pump is a one stage gear type pump and supplies oil underpressure to the engine bearings,gearbox drive and accessory drives. The oil ispumped through a pressure filter to remove any large debris.It has a cleanablefilter element.The pressure filter housing is installed at the oil pressure pump .The pressure filter housing incorporates a pressure priming connection and aantidrain valve to prevent oil loss during removal.

The filter does not have a bypass.The pressure pump housing incorporates the pressure filter ,a cold start pres-sure relief valve and a pressure pump flow trimming valve.The pressure relief valve bypasses the pressure circuit during cold starts.

LOCATIONThe pump is attached to the front face of the external gearbox on the left handside,just below the oil tank.

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79−00

Page: Page: 107FRA US/T bu August 2001 Page: 107Figure 54 Pressure Pump & Filter

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ENGINEOIL SYSTEM

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79−00


AIR COOLED OIL COOLER (ACOC)

LocationThe ACOC is mounted on the engine fan case.

OperationThe ACOC is a additional oil cooler which removes heat from the engine lubri-cating oil using fan air and maintains the oil temperature within the specifiedrange.The filtered oil flows through the air cooled oil cooler before being cooled againthrough the fuel cooled oil cooler.The cooling air and the oil flows through the air / oil heat exchanger are shownbelow.

Features oil bypass valve ACOC oil temperature thermocouple ( for heat management system ) modulated air flow as commanded by EEC ( heat management system ).

air flow regulated by air modulating valve. Fuel pressure operated actuator Feedback LVDT

ACOC AIR MODULATING VALVE FAIL SAFE POSITION : ”OPEN”

ACOC OIL TEMPERATURE THERMOCOUPLE(refer to 73-20 Heat Management System)The ACOC thermocouple is used for the heat management system which iscontrolled by the EEC.

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ACOC OIL TEMPERATURETHERMOCOUPLE

Page: 109Figure 55 ACOC Air Flow

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79−00


FUEL COOLED OIL COOLER (FCOC)

LocationThe oil passed through the ACOC flows through the Fuel Cooled Oil Cooler(FCOC) ,installed on the left hand side of the fan casing,before it is sent to thebearing compartments and both the angle and main gearboxes.

Purpose The FCOC cools the oil by using low pressure fuel. The FCOC also warms the low temperature fuel to the de-icing level. The FCOC has 2 bypass valves.

DescriptionThe FCOC consits of a housing containing a removable core,a header and afuel filter cap.The core is composed of vacuum brazed tubes through whichfuel passes.

Bypass valves One is an oil pressure relief bypass valve which diverts the excessive oil

pressure during engine cold start. The other is a fuel filter bypass valve which ensures fuel flow in the event of

fuel filter clogging.

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A

LOCATION

IN

OUT

DRAIN HOLE

OIL

Page: 111Figure 56 Fuel Cooled Oil Cooler

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SCAVENGE SYSTEMThe scavenge system main components are:

− chip detectors,− 6 scavenge pumps with strainers,− one common scavenge filter.− a 2−positions scavenge valve.( Bearing No.4 )

SCAVENGE PUMPS

PurposeThe scavenge pump returns the oil back to the oil tank.

DescriptionThe scavenge pump is a five−stage gear type pump on the rear left side of thegeabox.Four stages of the scavenge pump are two−gear displacement pumps .The stage used for the two main gearbox scavenge lines consists of threemeshing gears producing two inlets and outlets on opposite sides.All 6 scav-enge pumps are housed together as a single unit.The pump capacity is deter-mined by the width of the gears.

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SCAVENGE

Page: 113Figure 57 Scavenge Pump Assembly

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SCAVENGE OIL COMPONENTS

Scavenge FilterThe flows from the 6 scavenge pumps are mixed together at the scavenge filtercommon filter inlet.

LocationThe filter is mounted to the rear of the oil tank.

Features disposable filter element by-pass valve (opens when filter clogs) Differential pressure connections provides housing for the master magnetic chip detector Oil Temperature sensor

Scavenge Filter Differential Press. SwitchThe scavenge filter differential pressure switch is installed on a bracket at thetop left side of the engine fan case,near the FCOC.Switches the ECAM OIL FILTER CLOG warning when the filter becomesblocked ( +12PSI or - 2 PSI differential press)

Engine Oil TemperatureThe scavenge oil temperature thermocouple is located in the combined scav-enge line between the master magnetic chip detector and the scavenge filterfor indication in the cockpit.The oil temperature is sensed by a dual resistor unit. The unit consists of asealed, wire−wound resistance element. This element causes a linear changein the DC resistance when exposed to a temperature change.Temperature measurement range:− 60 deg. C to 250 deg. C.The analog signal from the scavenge oil temperature thermocouple is trans-mitted to the EIU. The EIU transforms this signal into a digital signal.This digitalsignal is then transmitted to the lower ECAM display unit through the FWCsand the DMC.

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79−00


SCAVENGE OIL FILTER

ELECTRICAL CONNECTOR

SEAL - RING

OIL TEMP. SENSOR

OIL TEMP. SENSOR

SCAVENGE FILTER DIFFERENTIAL(PRESS. DROP.) WARNINGSWITCH ( DELTA P. 12 PSI )

Page: 115Figure 58 Scavenge Filter,Delta P.Sw and Oil Temp. Sensor

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DE-OILER

LocationThe de-oiler is bolted to the right hand front face of the external gearbox.

Purpose To separate the breather air/oil mixture. return the oil to the oil scavenge system via its own scavenge pump. vent the air overboard through the R/H fan cowl.

Features provides mounting for the No.4 bearing chamber scavenge valve. overboard vent. provides location for the No.4 bearing magnetic chip detector housing.

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FROM NO 4 BEARINGSCAVENGE VALVE.

BREATER AIR

FROM OIL TANK

Page: 117Figure 59 De-Oiler

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NO4 BEARING SCAVENGE VALVE

LocationThe valve is mounted on the front face of the de-oiler casing.

PurposeMaintains No.4 bearing compartment seal differential pressure to reduce over-board loss of vent air and to prevent deteriation of the carbon seals by restrict-ing the venting of the compartment air/oil mixture to the de-oiler.

Type of valvePneumatically operated two position valve.

Features Position feed back signal to EIU ( reed switch ) uses stage 10 air as servo air uses value of pressure of stage 10 air as operating parameter. Fully open at low engine speeds( stage 10 air less than 150 PSI ) Minimum open at high engine speed (stage 10 air more than 200 PSI )

NO 4 BEARING PRESSURE TRANSDUCER

PurposeThe purpose of the No.4 bearing indicating system is to monitor the correct operation of the No.4 bearing 2−position scavenge valve and to detect a No.4 bearing carbon−seal failure.The No.4 bearing pressure transducer is installed on the right side of the deoiler and senses pressure at the No.4 bearing outlet line.Linear output 1VDC to 9 VDC (0 To 300 PSIG),

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79−00


NO.4 BEARING PRESSURETRANSDUCER

A

A

NO.4 BEARING OIL INLET

POSITION REED SWITCH

10TH STAGE AIR

DE-OILER CASE


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NO4 BEAR. SCAV. VALVE DESCRIPTIONOperationThere are two basic operating positions, low power and high power. In the low−power position, where the compressor 10th stage pressure (PS10)is less than 150 PSI, the valve is held spring loaded in the fully open position.The bearing compartment scavenge flow passes through the valve, restrictedonly by the porting in the valve seat.As the engine power increases, the PS10 pressure rises. When this pressureexceeds 150 psi, the valve moves away from the max flow stop. This is due tothe pressure acting on the differential areas of the valve and overcoming thespring load. The valve moves towards the min flow or high power setting. Asthe valve moves towards the peripheral ports in the seat, totally closing these ports, the flow through the valve is now restricted to one central port in thevalve seat. Full travel is achieved at PS10 pressure of approximately 210 psi.As the valve moves away from the max flow stop, the influence of the magnetson the reed switch decreases and the reed switch opens.The circuit is broken, indicating that the valve has moved.As the engine power decreases, the spring load overcomes the decreasingPS10 pressure. The valve moves towards the max flow or low power position,uncovering the ports in the valve seat and restoring maximum flow through thevalve. As the valve approaches the maximum flow stop, the influence of themagnets on the reed switch increases.The reed switch closes, completing the circuit and indicating the valve position.

NO.4 BEARING SCAVENGE VALVE INDICATINGThe EIU incorporates three logics allowing the monitoring of the scav-enge valve operation as well as a No.4 bearing carbon - seal failure

LOW POWER SETTING:At engine low power, the bearing scavenge valve is open and the reedswitch on the valve closes providing a ground signal for the EIU logic.

HIGH POWER SETTING:At engine high power, the bearing scavenge valve closes (to maintain theNo.4 bearing pressure ratio in the bearing compartment) and the reedswitch on the valve opens.

The No.4 bearing internal pressure is measured by the No.4 bearing pressureXMTR in the oil return line to the deoiler.The transducer supplies a pressuresignal to one of the three EIU logics.

Two EIU logics provide a warning message to the ECAM :Eng.1 (2) BEARING 4 OIL SYS. ( class 2 )and a message SCAVENGE VALVE FAULT is displayed when the valve isnot in the correct position according to the sensed burner pressure.The massage HI PRESS is displayed when the No. 4 bearing compartmentpressure is is to high according to the valve position and a high burnerpress.(possible Carbon seal failure ) or scavenge valve stuck in closed orscavenge line pressure sensor malfunction.

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� � � � ��

LOW OIL PRESS.SWITCH

OIL PRESS XMTR

NO.4 BEARING PRESSXMTR

EIUEEC

MIN FLOWMAX FLOW

NO.4 BEARING SCAVENGEVALVE

OIL PRESSURE

FAN AIRBUFFER AIR COOLER

( ACAC)

TO OTHER BLEED SOLENOID VALVES

EEC

TODEOILER

PB

10THSTAGESOLENOIDVALVE

REED SW

OIL AND AIR

10 TH STAGE AIR (4X)

BEARING 4COMPARTMENT

STAGE 10 AIR

BUFFERAIR

MAKE UP AIR

MAKE UPAIR VALVE

STAGE 12COMBUSTIONCHAMBER


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ENGINE OIL PRESSUREThe Oil pressure is directly linked to the opening and closing of the No.4 Bear-ing Scavenge Valve.A closing of the valve (at approx. 85% N2 ) will restrict the return scavengeflow to the deoiler.This will result in a pressure drop,because the ratio of the pressures willchange. ( the oil pressure is the differential pressure of the oil pressure feedline and the scavenge line).The No. 4 compartment scavenge oil pressure range is 0 to 160 PSI . Normal operating pressure is 0-145 PSI after three minutes of stabilization atidle speed.

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Page: Page: 123FRA US/T bu August 2001 Page: 123Figure 62 Oil Pressure Chart

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OIL SYSTEM PRESSURE SENSING

GeneralThe oil pressure indicating system gives a cockpit indication of theengine oil system working pressure.The indication of this pressure comes electrically from an oil pressure transmitter on each engine. The oil pressure transmitter is bolted to a bracket on the top left side of the

engine fan case. The oil pressure transmitter is connected to the engine oil system by

two steel tubes. One tube connects to the oil supply tube (to the engineand gearbox bearings). The other tube connects to the No. 4 bearing oilscavenge tube (to the oil scavenge pump).

Power supply : 28VDC from busbar 101PP (202PP). Pressure range : 0 to 400 psid. Output voltage : 1VDC to 9VDC varying linearly with pressure from 0 to

400 psid.

LOW OIL PRESSURE SWITCHThe low oil pressure switch is installed on a bracket at the top left side of theengine fan case,beside the oil pressure transmitter.The oil pressure switch is connected between the oil supply tube and the No.4bearing scavenge tube.When the oil pressure drops below 60 psi the switch closes and a red warningis triggert in the cockpit.The set point range is between 45psi and 75psi.

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Low Oil Press Switch

Pressure PortScavenge Oil Pressure Port

LOCATION

Oil Press. Transmitter

Page: 125Figure 63 LOP Switch and Oil Press. Transmitter

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MAGNETIC CHIP DETECTORS (M.C.D.)

A total of 7 M.C.D. ‘s are used in the oil scavenge system.Each bearing compartment and gearbox has its own deticated M.C.D. (two inthe case of the main gearbox)although that for the No.4 bearing is located inthe de-oiler scavenge outlet).

Magnetic Chip Detectors LocationThe M.C.D. ‘s for: No.1,2 and 3 bearings main gearbox / L/H scavenge pick-up angle gearbox

are located to the rear of the main gearbox on the L/H side ,as shown below.

The M.C.D.‘s for: No.5 bearing De - oiler ( No.4 bearing ) Main gearbox ( R/H scavenge pick up )

are located as shown below.

CAUTION: DO NOT TRY TO INSTALL THE MCD IF THE SEAL RINGSARE NOT INSTALLED.A SAFTEY MECHANISM ISINSTALLED IN THE MCD HOUSING TO PREVENT INSTAL-LATION OF THE MCD IF THE FRONT SEAL RING IS NOTINSTALLED.IF ONLY THE FRONT SEAL RING IS INSTALLED , FAILUREOF THIS SEAL RING COULD RESULT IN AN IN-FLIGHTSHUTDOWN OF THE ENGINE BECAUSE OF OIL LEAKAGE.

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No. 4 BEARING

Page: 127Figure 64 Chip Detectors

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MASTER MAGNETIC CHIP DETECTORThe master chip detector is located in the combined scavenge return linie,onthe scavenge filter housing.The Master Chip Detector is accessible through its own access panel in theL/H fan cowl.If the master M.C.D. indicates a problem then each of the other M.C.D.‘s isinspected to indicate the source of the problem.

CAUTION: DO NOT TRY TO INSTALL THE MCD IF THE SEAL RINGSARE NOT INSTALLED.A SAFTEY MECHANISM ISINSTALLED IN THE MCD HOUSING TO PREVENT INSTAL-LATION OF THE MCD IF THE FRONT SEAL RING IS NOTINSTALLED.IF ONLY THE FRONT SEAL RING IS INSTALLED , FAILUREOF THIS SEAL RING COULD RESULT IN AN IN-FLIGHTSHUTDOWN OF THE ENGINE BECAUSE OF OIL LEAKAGE.

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ENGINEOIL SYSTEM

A319/A320/A321IAE V2530-A5

79−00

Page: Page: 129FRA US/T bu August 2001 Page: 129Figure 65 Master Magnetic Chip Detector

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ENGINEOIL SYSTEM

A319/A320/A321IAE V2530-A5

79−00


IDG OIL SERVICING

IDG oil pressure fillA quick fill coupling situated on the transmission casing enables pressure fillingor topping up the unit with oil. The oil thus introduced flows to the transmissionvia the scavenge filter and external cooler circuit. This ensures : − the priming of the external circuit − the filtration of any oil introduced.An internal standpipe connected to an overflow drain ensures a correct quantityof oil.

Oil filterA clogged filter indication is provided by a local visual pop out indicator. Theindicator is installed on the anti drive end of the IDG.

Oil level checkYou can read the oil level through two sight glasses located on the IDG. One sight glass serves for the CFM 56 engine, the other one for the V2500engine. The oil level must be at or near the linie between the yellow and green

bands. If the oil level is not at this position,connect the overflow drain hose and

drain the oil until the correct filling level is reached.This will also depressu-rize the IDG case.

NOTE: IF THE OVERFLOW DRAINAGE PROCEDURE IS USED IT CANTAKE UP TO 20 MINUTES TO COMPLETE.FAILURE TO OBSERVE THE OVERFLOW TIME REQUIREMENTSCAN CAUSE HIGH OIL LEVEL CONDITION RESULTING IN ELE-VATED OPERATING TEMPERATURES AND DAMAGE/DISCON-NECT TO IDG.

Page: 130

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ENGINEOIL SYSTEM

A319/A320/A321IAE V2530-A5

79−00


Press Fill Valve

Dust CapDust Cap

Overflow DrainValve

V2500

A

Page: 131Figure 66 IDG Oil Servicing

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ENGINE FUEL AND CONTROLGENERAL

A319/A320/A321IAE V2530−A5

73−00



73−00 FUEL SYSTEM PRESENTATIONGeneralThe fuel system enables the combustion of fuel under appropriate conditions offlow rate and pressure. The FADEC controls the fuel supply via the Fuel Meter-ing Unit (FMU). High pressure fuel is also used to provide pressure for someactuators.The major components are − High and low pressure fuel pumps (dual unit) − Fueloil heat exchanger − Low pressure fuel filter − Fuel Metering Unit (FMU) − Fuel distribution valve − 20 fuel injectors − Diverter and return to tank valve − IDG fueloil heat exchanger.

DistributionThe fuel supplied from aircraft tanks flows through a centrifugal pump (LPstage) then through the Fuel Cooled Oil Cooler and then through a filter anda gear pump (HP stage).The fuel from the HP pump is delivered to the Fuel Metering Unit (FMU) which controls the fuel flow supplied to the fuel nozzles (through the fuel flow meterand the fuel distribution valve).The FMU also provides hydraulic pressure to all hydraulic system external actuators. These include the Booster Stage Bleed Valve actuators, Stator VaneActuator, ACOC air modulating valve and HPT/LPT Active Clearance Controlvalve. Low pressure return fuel from the actuators is routed back into the fueldiverter valve.The fuel diverter and return to tank valve enables the selection of four basicconfigurations between which the flow paths of the fuel in the engine are variedto maintain the critical IDG oil, engine oil and fuel temperatures within speci-fied limits.The transfer between configurations is determined by a softwarelogic contained in the EEC.

ControllingThe Fuel Authority Digital Electronic Control (FADEC) system provides fullrange control of the engine to achieve steady state and transient performance when operated in combination with aircraft subsystems.The FADEC is a dualchannel EEC with crosstalk and failure detection capability.In case of specificfailure detection, the FADEC switches from one channel to the other.

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A319/A320/A321IAE V2530−A5

73−00

Page: 133FRA US/T Bu August 2001 Page: 133Figure 67 Fuel System Schematic

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A319/A320/A321IAE V2530−A5

73−00


73−00 FUEL SYSTEM PRESENTATION

GeneralThe fuel system enables delivery of a fuel flow corresponding to the power required and compatible with engine limits.The system consists of:

− the two stage fuel pump with low pressure & high pressure ele ments,− the engine fuel cooled oil cooler (FCOC),− the fuel filter− the fuel diverter and return to tank valve.− the integrated drive generator (IDG) fuel cooled oil cooler (FCOC),− the fuel metering unit (FMU),− the fuel distribution valve,− the fuel flow transmitter,− 20 fuel nozzles,

DESCRIPTION AND OPERATION

DistributionThe fuel supplied from aircraft tanks flows through a centrifugal pump (LPstage) then through the Fuel Cooled Oil Cooler and then through a filter anda gear pump (HP stage).The fuel from the HP pump is delivered to the Fuel Metering Unit (FMU) which controls the fuel flow supplied to the fuel nozzles (through the fuel flow meterand the fuel distribution valve).The FMU also provides hydraulic pressure to all hydraulic system external actuators. These include the Booster Stage Bleed Valve actuators, Stator VaneActuator, ACOC air modulating valve and HPT/LPT Active Clearance Controlvalve. Low pressure return fuel from the actuators is routed back into the fueldiverter valve.The fuel diverter and return to tank valve enables the selection of four basicconfigurations between which the flow paths of the fuel in the engine are variedto maintain the critical IDG oil, engine oil and fuel temperatures within speci-

fied limits.The transfer between configurations is determined by a softwarelogic contained in the EEC.

ControllingThe Fuel Authority Digital Electronic Control (FADEC) system provides fullrange control of the engine to achieve steady state and transient performance when operated in combination with aircraft subsystems.The FADEC is a dualchannel EEC with crosstalk and failure detection capability.In case of specificfailure detection, the FADEC switches from one channel to the other.The FADEC System operates compatibly with applicable aircraft systemsto perform the following:

− Control of fuel flow, stator vanes and bleeds to automatically maintain for ward and reverse thrust settings and to provide satisfactory transient response.− Protect the powerplant from exceeding limits for N1, N2, maximum allow able thrust, and burner pressure.− Control of the low and high turbine active clearance control systems.− Control of fuel, engine and IDG oil temperature.− Control of the thrust reverser.− Automatic sequencing of start system components.− Extensive diagnostic and maintenance capability.

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A319/A320/A321IAE V2530−A5

73−00


TANKFUELTEMPSNSR

RVDT

SDAC

DMC FWC

Page: 135Figure 68 Fuel System Schematic

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ENGINE FUEL AND CONTROLINDICATING

A319/A320/A321IAE V2530−A5

73−30


ATA 73−30 INDICATINGGENERAL

IndicatingThe engine fuel system is monitored from:− the ECAM display,− the warning and caution lights.The indications cover all the main engine parameters through the FADEC.The warning and cautions reflect:− the engine health and status through the FADEC,− the FADEC health & status,− the fuel filter condition through a dedicated hardwired pressure switch.The fuel system is monitored by: The fuel flow indication on the upper ECAM display unit permanently

displayed in green and under numerical form. The fuel filter clogging caution (amber) on the lower ECAM display unit

associated with the MASTER CAUT light and the aural warning (single-chime).

Fuel flow indication, Fuel Used

The Fuel Flow Transmitter is installed near the FMU. The signals are routed tothe EEC and via the DMCs to the ECAM.

The Fuel Used-is calculated in the DMCs .

The fuel flow transmitter signal is fed to the FADEC which processes it andtransmits the information to the ECAM system for display .

Fuel filter clogging indicationThe fuel filter clog indication is provided on the lower ECAM display unit. Whenthe pressure loss in the fuel filter exceeds 5 plus or minus 2 psid, the pressureswitch is energized.This causes:

− Triggering of the MASTER CAUTion light and single chime.− The engine page to come on the lower ECAM DU with the caution signal FUEL CLOG.− The associated caution message to come on the upper ECAM DU.

When the pressure loss in the filter decreases between 0 and −1.5 psid fromthe filter clog energizing pressure, the pressure switch is de−energized whichcauses the caution to go off.The differential pressure switch signal is fed directly to the SDACthrough the hardware .

Page: 136

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ENGINE FUEL AND CONTROLINDICATING

A319/A320/A321IAE V2530−A5

73−30


� � � � � � � � � ��

� � � � � � � � � ��

� � � � � ��

� � � � � ��

� � � � � � � ��

KG/H2500 2500

13000 KG

Page: 137Figure 69 Fuel System Indication

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ENGINE FUEL AND CONTROLCONTROLLING

A319/A320/A321IAE V2530 A5

73−20


FUEL PUMP

GeneralThe LP / HP fuel pumps are housed in a single pump unit which is driven by a common gearbox output shaft. A low pressure (LP) stage and a highpressure ( HP ) stage provide fuel at the flows and pressures required for operation of hydromechanical components and for combustion in the burner.The unit consists of a LP centrifugal boost stage which feeds an HPsingle stage, two gear pump. The housing has provision for mounting the fuel metering unit ( FMU ).The LP stage receives fuel from aircraft tanks through the aircraft pumps.The LP pump is designed to provide fuel to the HP gear stage with theaircraft pumps inoperative. After passing through the LP boost stage, fuelproceeds through the fuel filter to the HP gear stage. A coarse meshstrainer is provided at the inlet to the HP gear stage. This stage is protectedfrom overpressure by a relief valve. Exceeding flow from the gearstage pump isrecirculated through the FMU bypass loop to the low pressure side of thepump.

FUEL METERING UNITThe FMU is the interface between the EEC and the fuel system.It is located on the dual fuel pumps unit, on the rear of the main gearbox, andis retained by four bolts as shown below.All the fuel delivered by the HP fuel pumps - which is much more than the engine requires - passes to the F.M.U. The FMU, under the control of theEEC meters the fuel supply to the spray nozzles. It also supplies HP fuel forthe operation ( muscle ) of a number of actuators. Any fuel supplied by theHP pumps which is not needed for these two uses is returned, from the FMU tothe LP side of the fuel system.In addition to the fuel metering function the FMU also houses the : Overspeed Valve Pressure Raising and Shut Off Valve

The overspeed valve under the control of the EEC, provides overspeed protec-tion for the LP ( N1 ) and HP ( N2 ) rotors.The Pressure Raising and Shut Off Valve provides isolation of the fuel suppliesat engine stop .

NOTE: THERE ARE NO MECHANICAL INPUTS TO, OR OUTPUTS FROMTHE FMU.

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A319/A320/A321IAE V2530 A5

73−20

Page: 139FRA US/T bu August 2001 Page: 139Figure 70 Fuel Pump and Fuel Metering Unit

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ENGINE FUEL AND CONTROLDISTRIBUTION

A319/A320/A321IAE V2530−5A

73−10


ATA 73−10 FUEL DISTRIBUTION COMPONENTSFUEL FILTER

DescriptionThe fuel filter element is a low pressure filter which removes all contaminationfrom fuel to go through it.The filter element is installed in the lower housing of a fuel cooled oil cooler ( FCOC ). The FCOC includes the following components :

– A filter cap which has a pressure plate to keep the filter element in position once installed.The filter cap of the FCOC also includes a fuel drain plug to drain the fuel for maintenance purposes.

– A filter bypass valve to let the fuel go around the filter element when it be comes clogged.

FUEL FILTER DIFF. PRESS. SWITCH

The fuel filter clog indication is provided on the lower ECAM display unit. Whenthe pressure loss in the fuel filter exceeds 5 plus or minus 2 psid, the pressureswitch is energized.When the pressure loss in the filter decreases between 0 and −1.5 psid fromthe filter clog energizing pressure, the pressure switch is de − energized which causes the caution to go off.The differential pressure switch signal is fed directly to the SDAC

FUEL TEMPERATURE THERMOCOUPLE( refer to 73-20 Heat Management System )The measured temperature is transmitted to the EEC ( Electronic EngineControl ) and used for the Heat Management System.


GeneralThe fuel diverter and return valve ( FD & RV ) is a primary unit in the heat management system ( HMS ) of the engine. The FD & RV has two valves inone body. They are a fuel diverter valve (FDV) and a fuel return valve ( FRV ).The FDV operates to change the direction of the fuel metering unit ( FMU )spill flow to :

− The fuel cooled oil cooler ( FCOC ) or,− the fuel filter ( element ) inlet or,− the fuel cooled IDG oil cooler ( IDG FCOC ).

The FRV operates to control fuel flow which goes back to the aircraft fuel tankacting as a fuel cooler.

Page: 140

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A319/A320/A321IAE V2530−5A

73−10


FCOC

A

LOW PRESS FUEL FILTER

FUEL COOLEDOIL COOLER ( FCOC )

FUEL DIVERTERAND RETURN VALVE ( FDRV )

FCOC FUEL TEMP. THERMOCOUPLE

FUEL FILTER DIFFERENTIALPRESSURE SWITCH

A

CONNECTION TO AIRCRAFT FUEL TANK

FCOC INLET

FUEL FILTER DIFF.PRESS. SW.

Page: 141Figure 71 Fuel Filter Diff. Press. Switch/FCOC Fuel Temp. Thermocouple

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A319/A320/A321IAE V2530−5A

73−10


FUEL DISTRIBUTION VALVEGeneralThe fuel distribution valve ( FDV ) subdivides scheduled engine fuel flowfrom the fuel metering unit (FMU) equally to ten fuel manifolds, each ofwhich in turn feeds two nozzles.

DescriptionThe fuel distribution valve is installed at the 4:00 o’clock location, at the frontflange of the diffuser case.The fuel distribution valve receives fuel through a fuel line from the fuel meter-ing unit. The fuel goes through a 200 micron strainer, and then into ten internaldischarge ports. The ten discharge ports are connected to the ten fuel manifolds.Eight of the ten internal discharge ports in the valve are connected afteran engine shutdown. Eight of the fuel manifolds are drained into the engine through the lowest fuelnozzle.The two fuel manifolds which remain full help supply fuel for the nextengine start.

FUEL NOZZLES

FUEL DISTRIBUTIONVALVE

HP/LP PUMPS

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A319/A320/A321IAE V2530−5A

73−10

Page: 143FRA US/T Bu August 2001 Page: 143Figure 72 Fuel Distribution Valve

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A319/A320/A321IAE V2530−5A

73−10


FUEL MANIFOLD AND TUBES

DescriptionThe fuel manifold and fuel tubes consist of several single wall tubes which carry fuel between components in the fuel system. Fuel supplied to the fuelnozzles is carried by a large tube from the fuel metering unit to the fuel distribu-tion valve. At the fuel distribution valve the fuel supply is split and carried totwenty fuel nozzles by ten manifolds.Each fuel manifold feeds two fuel nozzles. Fuel pressure for actuating variousvalves is supplied by small tubes from the fuel metering unit mounted on thefuel pump.All the brackets and tubings are fire proof.

FUEL NOZZLE

GeneralThe fuel nozzles receive fuel from the fuel manifolds. The fuel nozzles mix thefuel with air, and send the mixture into the combustion chamber in a controlledpattern.

Description/OperationThere are 20 fuel nozzles equally spaced around the diffuser case assembly.The fuel nozzles are installed through the wall of the case, and each nozzle isheld in position by three bolts.The fuel nozzles carry the fuel through a single orifice. The fuel is vaporized byhigh−velocity air as it enters the combustion chamber. The fuel nozzle formsthe atomized mixture of fuel and air into the correct pattern for satisfactorycombustion.The design of the fuel nozzle results in fast vaporization of the fuel through thefull range of operation. This results in decreased emissions, high combustionefficiency, and good start quality.The high−velocity flow of fuel prevents formation of coke on areas where fueltouches metal. Heatshields installed also prevent formation of coke.

Page: 144

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A319/A320/A321IAE V2530−5A

73−10

Page: 145FRA US/T Bu August 2001 Page: 145Figure 73 Fuel Distribution Tubes

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A319/A320/A321IAE V2530−5A

73−10


IDG FUEL COOLED OIL COOLERThe IDG oil cooler is installed at the left hand side on the fan case, near theFCOC.The IDG oil cooler has two sets of inlet and outlet ports. One set of ports isused for the flow of the fuel to or from the fuel diverter and return valve. Theother set of ports is used for the flow of oil from and to the IDG.The hot scavenge oil which has been used to lubricate and cool the IDG, flowsfrom the IDG to the oil cooler.As the oil goes through the oil cooler, the heat in the oil is transmitted to thefuel. The cooled oil then returns to the IDG.Two drain plugs are also installed in the oil cooler, one for the fuel and one forthe oil.

IDG OIL COOLER TEMP. THERMOCOUPLE( refer to 73-20 Heat Management system )This temperature information is send to the EEC and is used for the heat management system.

Page: 146

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A319/A320/A321IAE V2530−5A

73−10


IDG FUEL COOLED OILCOOLER

IDG OIL TEMP.THERMOCOUPLE

FUEL INLET / OUTLET

OIL INLET

OIL OUTLET

DRAIN PLUGS

Page: 147Figure 74 IDG Fuel Cooled Oil Cooler

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A319/A320/A321IAE V2530 A5

73−20


FUEL METERING UNIT

GeneralA simplified schematic representation of the Fuel Metering Unit is shownbelow.The three main functions of the FMU are : metering the fuel supplies to the fuel spray nozzles. overspeed protection for both the LP ( N1 ) and HP ( N2 ) rotors. isolation of fuel supplies for starting/ stopping the engine.

These three functions are carried out by three valves arranged in series, asshown: the Fuel Metering Valve the Overspeed Valve the Pressure Raising and Shut Off Valve.

The position of each valve is monitored and positional information is trans-mitted back to the EEC.This ensures that the EEC always knows that the valves are in the commandedposition.

FAIL SAFE POSITION OF THE METERING VALVE TORQUE MOTOR :” MINIMUM FUEL FLOW CONDITION ”

Overspeed Valve

OperationThe overspeed valve is spring loaded to the closed position, it is opened byincreasing fuel pressure during engine start and during normal engine operationis always fully open.In the event of an overspeed ( 109,1% N1 , 105,4% N2 ) the EEC sendsa signal to the overspeed valve torque motor which changes position and di-rects H.P. fuel to the top of the overspeed valve − this fully closing the valve.A small by − pass flow is arranged around the overspeed valve to prevent engine flame out.

The overspeed valve is hydraulically latched in the closed position, thus pre-venting the engine from being reaccelerated.The recommended procedure is for the flight crew to shut down the engine .To shut down the engine is the only way to release the hydraulic latching.

NOTE: BECAUSE THE OVERSPEED VALVE IS SPRING LOADED TOTHE CLOSED POSITION, AND OPENED BY FUEL PRESSURE,THE OVERSPEED VALVE WILL CLOSE ON EVERY ENGINESHUT DOWN.

FAIL SAFE POSITION: ” NORMAL FUEL METERING”

Pressure Raising and Shut off Valve

The PRSOV torque motor is commanded open by the EEC during AUTO startsor Closed by the EEC during AUTO start sequences if the sequence has to bestopped for any reason.It is commanded open or closed by the MASTER SWITCH in the cockpit duringMANUAL starts.

NOTE: THE EEC’S ABILITY TO CLOSE THE SHUT OFF VALVE IS INHIB-ITED ABOVE 43% N2. ABOVE 43% N2, AND IN FLIGHT, THEPRSOV CAN ONLY BE CLOSED BY THEMASTER SWITCH IN THE COCKPIT.

FAIL SAFE POSITION OF THE PRSOV : ” LAST COMMANDED POSITION ”

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A319/A320/A321IAE V2530 A5

73−20


VARIABLE 2 POS. 2 POS.

MASTER

LEVER

Page: 149Figure 75 Fuel Metering Unit Schematic

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A319/A320/A321V2530-A5

73-10


HP & LP FUEL SOV CONTROL

The HP fuel shut off valve control is fully electrical. It is performed from the engine panel in the cockpit as follows :

Opening of the HP fuel PRSOV :It is controlled by the EEC : the EEC receives the commands from theMASTER control switch and ignition selector switch.

Closure of the HP fuel PRSOV :It is controlled directly from the MASTER control switch in OFF position

PRSOV Fuel Shut Off ControlThe FADEC control system contains a fuel shut − off in the FMU , which actsthrough a 2 position torque motor to close the pressurizing valve :The fuel shut − off is direct−hardwired to the MASTER control switch.This tourque motor operated PRSOV is powered by the 28VDC. Loss of power supply does not lead to change the selected HP fuel

shutoff valve position. The cockpit command ” OFF ” has priority over the EEC command.

LP Fuel Shutoff Valve ControlThe LP fuel shut−off system has two independent electrical control circuitsfor each LP fuel − valve. They connect through a control relay to these related switches :− the ENG MASTER switch− the FIRE PUSH switch .When the No. 1 ENG MASTER switch is set to ON, it disconnects a 28VDCsupply from the relay 11QG ( HP FUEL SOV SOL P / B SW ). The relay11QG de − energizes and connects a 28VDC supply ( through the ENG 1FIRE PUSH switch ) to the ” open ” side of the LP fuel − valve actuator. The actuator then opens the LP fuel − valve.When the No. 1 ENG MASTER switch is set to OFF, it connects a 28VDC sup-ply to the relay 11QG. The relay energizes and connects a 28VDC supply ( through the ENG 1 FIRE PUSH switch ) to the ” close ” side LP fuel − valveactuator. The actuator then closes the LP fuel − valve.If the ENG 1 FIRE PUSH switch is operated :

− it disconnects the 28VDC supply to the ” open ” side of the LP fuel − valve actuator− it connects a 28VDC supply to the ” close ” side of the LP fuel valve actuator the LP fuel − valve moves to the closed position.

NOTE: THE LP FUEL − VALVE OPENS ( CLOSES ) WHEN THE ENGMASTER SWITCH IS SET TO ON ( OFF ). BUT THE OPERATIONOF THE ENGINE FIRE PUSH SWITCH ALWAYS OVERRIDES ANON SELECTION AND CLOSES THE VALVE.

NOTE: IT IS ALSO COMMANDED OPEN VIA THE RELAY 11QG WHENTHE C / B OF THE HP FUEL SOV IS PULLED, ( RELAY 11QG( 12QG ) DEENERGIZED ).

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A319/A320/A321V2530-A5

73-10


EECFMU

MASTER SW 1

CENTRAL PEDESTAL 115VU

HP FUELSOV

2 POS TM

LP FUEL SHUTOFF VALVE 1

HP FUELSOV

CLOSEDPOS SW‘s

ENGINE 1HP FUEL SOV

49VU A1

28 V DCESS

28 V ESS

28 V DC 2

121VU M25

ENGINE 1FUEL LPVALVE MOT 2

49VU A8

ENGINE 1FUEL LPVALVE MOT 1

M 1

M 2

TO ECAM

SHUT

OPEN

OPEN

SHUT

CLOSED

CLOSED

11QGRELAYENG / MASTER 1123VU 126

FIRE

FAULT

FIRE

FAULT

ENG115VU

MASTER 1 MASTER 2ON

OFFOFF MODENORM

CRANKIGN

START

1 2

ENG FIRE

PUSH1

ENG ENG1 2

VLV POSSW‘s

ON

Page: 151Figure 76 HP and LP Fuel Shutoff Valve ( SOV )

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ENGINE FUEL AND CONTROLFUEL DISTRIBUTION

A319/A320/A321V2530-A5

28−20


The engine fuel supply system has two fuel shut off valves. one PRSOV in the FMU One LP - fuel shut off valve on the front wing spar.

LOW PRESSURE FUEL SHUT OFF VALVEThe LP fuel − valve 12QM ( 13QM ) is in the fuel supply line to its relatedengine. The LP fuel − valve is usually open and in this configuration lets fuelthrough to its related engine. When one of the LP fuel − valves is closed, thefuel is isolated from that LP fuel valve’s related engine.The LP fuel − valve is installed between the engine pylon and the front face ofthe wing front spar ( between RIB 8 and RIB 9 ).Each LP valve has an actuator 9QG ( 10QG ). The interface between the actuator and the LP valve is a valve spindle. When the actuator is energized, itmoves the LP valve to the open or closed position. A V − band clamp80QM(81QM) attaches the actuator to the LP valve.Each actuator has two motors, which get their power supply from different sources :

− the 28VDC BATT BUS supplies the motor 1− the 28VDC BUS 2 supplies the motor 2.

If damage occurs to the electrical circuit, it is necessary to make sure that thevalve can still operate. Thus the electrical supply to each motor goes through adifferent routing. The routing for motor 1 is along the front spar. The routing for motor 2 is along the rear spar and then forward through the flap track fairing at RIB 6.The actuators send position data to the System Data − Aquisition Concentra-tors ( SDAC1 and SDAC2 ). The SDACs process the data and send it to theECAM which shows the information on the FUEL page.

Component DescriptionThe LP fuel − valve has:

− a valve body− a ball valve− a valve spindle− a mounting flange.

The LP fuel − valve actuator has two electrical motors which drive the same differential − gear to turn the ball valve through 90 deg. The limit switches inthe actuator control this 90 deg. movement and set the electrical circuit for thenext operation. One of the two motors can open or close the valve if the othermotor does not operate.The actuator drive shaft has a see/feel indicator where it goes through the actuator body. The see/feel indicator gives an indication of the valveposition without removal of the fuel LP fuel valve.

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ENGINE FUEL AND CONTROLFUEL DISTRIBUTION

A319/A320/A321V2530-A5

28−20


ELECTRICAL CONNECTORS

V-Clamp

Page: 153Figure 77 LP Fuel Shut−Off Valve

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ENGINE FUEL AND CONTROLHEAT MANAGEMENT SYSTEM

A319/A320/A321IAE V2530−A5

73−20


ATA 73-20 HEAT MANAGEMENT SYSTEMPRESENTATION

GeneralHeating and cooling of fuel, engine oil and IDG oil is accomplished by the FuelCooled Oil Cooler ( FCOC ), the Air Cooled Oil Cooler ( ACOC ) and the IDGcooler under the management of the EEC.

FUEL TEMPERATURE :The fuel temperature is measured at the exit of the filter.

OIL TEMPERATURTE :The engine oil temperature is measured upstream of the ACOC. The IDG oil temperature is measured at IDG oil cooler exit.The system is designed to provide adequate cooling, to maintain the critical oiland fuel temperatures within specified limits, whilst minimising the requirementfor fan air offtake.Three sources of cooling are available : the LP fuel passing to the engine fuel system the LP fuel which is returned to the aircraft fuel tanks fan air

There are four basic configurations between which the flow paths of fuel in theengine L.P. fuel system are varied.Within each configuration the cooling capacity may be varied by control valves which form the Fuel Diverter andBack to Tank Valve.The transfer between modes of operation is determined by software logic con-tained in the EEC. The logic is generated around the limiting temperatures ofthe fuel and oil within the system together with the signal from the aircraftwhich permits/inhibits fuel spill to aircraft tanks.

OperationThe measured temperature is transmitted to the EEC ( Electronic Engine Con-trol ). In response to the measured temperature, the EEC sends the signal tothe fuel diverter valve. The fuel diverter valve is used to reduce too high fueltemperature. The excess of high pressure fuel flow from the FMU ( FuelMetering Unit ) and return fuel from control actuator are plumbed to the di-verter valve which normally turns the flow to the FCOC exit.

FUEL TEMP. THERMOCOUPLEThe Fuel Temperature is measured by the thermocouple at the fuel exit of theFCOC ( Fuel Cooled Oil Cooler ).The thermocouple is composed of stainless steel sheathed sensing portion,stainless steel installing flange with seal spigot and electrical connector.The control of fuel temperature is done by the fuel diverter valve which isinstalled upstream of the FCOC.

IDG OIL COOLER TEMP. THERMOCOUPLEIDG Fuel Cooled Oil Cooler oil temperature is measured at the IDG Oil CoolerExit by a thermocouple.The termocouple gives an electrical output in relation to the temperature of theoil in the fuel cooled IDG oil cooler.This temperature information is send to the EEC and is used for the heat management system.

ACOC OIL TEMP. THERMOCOUPLE

The oil temperature is measured at the ACOC inlet by a thermocouple.Thethermocouple is composed of stainless steel sheathed sensing portion, stain-less steel installing flange with seal spigot and electrical connector.The temperature is transmitted to the EEC ( Electronic Engine Control ). Inresponse to the measured temperature, the EEC sends the signal to the modu-lating air valve.

ACOC MODULATING AIR VALVE

The modulating air valve regulates air flow to the ACOC. Oil heated by the en-gine passes through the ACOC and then to the FCOC. The air valve is modu-lated by the EEC to maintain both oil and fuel temperatures within acceptableminimum and maximum limits. Minimum oil temperature limits are used suchthat the oil may be used to prevent fuel icing with the use of FCOC. Maximumlimits have been established to avoid breakdown of engine oil and to avoid ex-cessively high fuel temperatures.

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73−20


IDG OIL TEMP.THERMOCOUPLE

FUEL TEMP. THERMOCOUPLE

OIL TEMP.THERMOCOUPLE

IDG OIL COOLER

FCOC

ACOC


EEC

Page: 155Figure 78 HMS Main System Components

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GeneralThe FDRV configuration allows four modes of operation according toelectrical signals from the EEC ( based on fuel and oil temperature measure-ments transmitted by thermocouples ).

DescriptionThe fuel diverter and return valve is installed on the FCOC.The FDV is a two − position selector valve which has two pistons in a sleeve.The two pistons are mechanically connected and make two valve areas whichare referred to as valve A and valve B. The FRV has a main valve and a push-ing piston in a sleeve. This main valve is a half − area piston − type valvewhich moves valve to change the metering port area. The main valve has twovalve functions that are referred to as valve C and valve D.The EEC gives the electrical signal to the FDRV to change the position of thevalves. The FDRV gives a feedback signal to the EEC to transmit the positionof valves in the unit. The fuel flow changes with the position of the valves.Thus, the fuel flow can be controlled through the FDRV and the EEC.

Fuel Return ValveThe EEC operates the dual−wound torque motor to control the servo pressure.This servo fuel pushes the main valve.The pressure balance between two sides of the main valve (Valves C and D)gives the direction and the speed of the valve movement.Then the valve changes the direction of the fuel flow and controls the meteringport area.

FAIL SAFE POSITION : ” FRV CLOSED, NO RETURN TO TANK ( MODE 3 or 5 )

Fuel Diverter ValveThe EEC energizes the solenoid valve to open the servo fuel flow. The switch assemblies transmit the EEC the valve position when the solenoidis de − energized.

FAIL SAFE POSITION :” FDV SOLENOID DE − ENERGIZED ” ( MODE 4 or 5 )

RETURN TO TANK MODES

HMS MODE 1 ( NORMAL MODE )This is the normal mode and is shown below. Fuel through the IDG FCOC orcombined with a quantity of fuel downstream of the FCOC is modulated for return to tank. FMU bypass flow is returned upstream of fuel filter. In this mode all the heat from the engine oil system and the I.D.G. oil system isabsorbed by the LP fuel flows. Some of the fuel is returned to the aircraft tankwhere the heat is absorbed or dissipated within the tank.

HMS MODE 4Fuel through IDG FCOC is modulated for fuel return to tank. FMUbypass flow returned upstream of FCOC. Supplemental cooling of fuelis provided by this mode.This mode is adopted at low engine speeds with a high IDG oil inlet temperature.In this mode the fuel / oil heat exchanger is operating as a fuel ” cooler ” and the heat passed to the engine oil is extracted by the air / oil heat exchanger.

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MODE 1 MODE 4Normal Return to Tank Mode Mode selected when in Normal

Mode 1 the Limit Temperature ( IDG Oil, Fuel ) can not be=maintained within Limits.

IDG FCOC ENG OILFCOC

FUEL RETURNTO TANK

VALVE

DIVERTERVALVE

FROM FUELTANK

TOINJECTORS

LP FUELSHUTOFFVALVE

LP PUMP

HP PUMP

FMU

RETURNTO

TANK

FUEL FILTER

OIL IN

OIL IN

OIL OUT

OIL OUT

FAN AIR

FUEL TEMP SNSR

OIL TEMPSNSR

OIL TEMPSNSR


FUEL RETURNTO TANK

VALVE

DIVERTERVALVE

FROM FUELTANK

TOINJECTORS

LP FUELSHUTOFFVALVE

LP PUMP

HP PUMP

FMU

RETURNTO

TANK

FUEL FILTER

OIL IN

OIL IN

OIL OUT

OIL OUT

FAN AIR

FUEL TEMP SNSR

OIL TEMPSNSR

OIL TEMPSNSR

ACOC ACOC

Page: 157Figure 79 Return to Tank Modes 1 and 4

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NO RETURN TO TANK MODES 3 AND 5

HMS MODE 3The third mode shown below is the mode adopted when the requirementsfor fuel spill back to tank can no longer be satisfied i. e.Fuel through IDG FCOC returned downstream of FCOC.FMU bypass flow returned upstream of fuel filter.Return to tank inhibited.This is the preferred mode of operation when return to tank is not allowed. Inthis condition all the heat from the engine and IDG oil systems is absorbed bythe burned fuel. If however, the fuel flow is too low to provide adequate cooling the engine oilwill be pre − cooled in the air/oil heat exchanger, by a modulated air flow, be-fore passing to the fuel / oil heat exchanger.

HMS MODE 5

Mode 5 is the mode which is used when system condition demand operation isas in Mode 3 but this mode is not permitted.FMU bypass flow returned upstream of FCOC via the IDG cooler in the reversedirection. Return to tank inhibited. This mode is adopted if the conditions exist.

NOTE: IN CASE THE OIL TEMPERATURE CANNOT BE KEPT WITHINTHE LIMITS THE FADEC SYSTEM WILL INCREASE THE ENGINESPEED ( FAIL SAFE POSITION ).

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MODE 3High Engine Speed


FUEL RETURNTO TANK

VALVE

DIVERTERVALVE

FROM FUELTANK

TOINJECTORS

LP FUELSHUTOFFVALVE

LP PUMP

HP PUMP

FMU

RETURNTO

TANK

FUEL FILTER

OIL IN

OIL OUT

OIL OUT

FAN AIR

FUEL TEMP SNSR

OIL TEMPSNSR

OIL TEMPSNSR


FUEL RETURNTO TANK

VALVE

DIVERTERVALVE

FROM FUELTANK

TOINJECTORS

LP FUELSHUTOFFVALVE

LP PUMP

HP PUMP

FMU

RETURNTO

TANK

FUEL FILTER

OIL IN

OIL IN

OIL OUT

OIL OUT

FAN AIR

FUEL TEMP SNSR

OIL TEMPSNSR

OIL TEMPSNSR

� Low Engine Speed Cold Fuel� Fail Safe Mode

MODE 5

OIL IN

ACOC ACOC

Page: 159Figure 80 NO Return to Tank Modes 3 and 5

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AIR MODULATING VALVE

PurposeTo govern the flow of cooling ( fan ) air through the air/oil heat exchanger( ACOC ), as commanded by the Heat Management Control System ( EEC )

TypePlate type supported at either end by stubshafts. operated by an Electro − Hydraulic Servo Valve mechanism.

LocationBolted to the outlet face of the air/oil heat exchanger.

Features fire seal forms an air tight seal between the unit outlet and the cowling ori-

fices controlled by either channel A or B of EEC valve positioned by fuel servo pressure acting on a control piston valve position feed back signal via LVDT to each channel of EEC fuel servo pressure directed by the Electro − Hydraulic Servo Valve

assembly which incorporates a Torque motor

FAIL SAVE POSITION :” AIR VALVE SPRING LOADED FULLY OPEN ” ( maximum cooling position)

In case of malfunction the warning ” ENG 1 ( 2 ) AIR EXCHANGER FAULT ” is displayed on the ECAM E / WD.

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Page: 161FRA US/T Bu August 2001 Page: 161Figure 81 Air Modulating Valve

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POWER PLANTDRAINS

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71-70


ATA 71-70 POWER PLANT DRAINS

GENERAL

The powerplant drain system collects fluids that may leak from some of the engine accessories and drives. The fluids collected from the power plant are discharged overboard through the drain mast installed below the engine acces-sory gearbox.The drain system comprises two sub−systems:

− fuel drains− oil, hydraulic and water drains

The two sub−systems come together at the same drain mast.

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POWER PLANTDRAINS

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71-70


ACOC

IDG

S. ( STARTER )

HYDRAULICS

FWD

LEFT SIDE

RIGHT SIDE

NOTE : CONNECTION * ARE AT THE ACCESSORY MOUNTING PAD ONLY

HYDRAULICPUMPS

AIRSTARTER

AIR COOLEDOIL COOLERACTUATOR

INTEGRATEDDRIVE

GENERATOR

OILTANK

SCUPPERFUEL

PUMPSFUEL

DIVERTERVALVE

FUELMETERING

UNIT

LP BOOSTERBLEED

MASTERACTUATOR

ACTUATOR

ACTUATOR

BIFURCATIONPANEL

VARIABLESTATORVANE

DRAINSMAST

BLEEDLP BOOSTER

SLAVE

ACTIVECLEARANCECONTROLACTUATOR

OILTANK

SCUPPER

OILTANK

SCUPPER

Page: 163Figure 82 Drain System

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POWER PLANTDRAINS

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71-70


PYLON DRAINSThe engine pylon is divided into 7 compartments.Various systems are routedthrough these areas.Any leckage from fluid lines is drained overboard through seperate lines in therear of the pylon.

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POWER PLANTDRAINS

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PYLON DRAINS

FUEL

FUEL / HYDR.

HYDR.

Page: 165Figure 83 Pylon Drains

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POWER PLANTDRAINS

A319/A320/A321IAE V2530-A5

71-70


DRAIN SYSTEM DESCRIPTION

Fuel DrainThe fuel drain lines come from engine accessories on the engine core, the engine fan case and gearbox. The engine core drains go through thebifurcation panel.The fuel drain system is connected to these engine accessories:

− Booster bleed master actuator (Core)− Booster bleed slave actuator (Core)− Variable Stator Vane Actuator (Core)− Active Clearance Control Actuator (Core)− Fuel diverter valve (FD)− Fuel metering unit (FMU)− LP/HP fuel pumps (FP)

Oil, Hydraulic and Water DrainsThe oil, hydraulic and water drains system comes from engine accessorieson the engine fan case and gearbox.The drain system is connected to these engine accessories:

− Air Cooled Oil Cooler actuator (ACOC) − Integrated Drive Generator (IDG) − Air starter (S) − Hydraulic Pump (HYD) − Oil tank scupper –Oil tank

The only hydraulic fluid drain is from the hydraulic pump. The other drains arefor engine oil or accessory lubricant.

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POWER PLANTDRAINS

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71-70

Page: 167FRA US/T Bu August 2001 Page: 167Figure 84 Drain System Leakage Test & Limits

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Engine ControlsGeneral

A319/A320/A321V2530-A5

76−00

Page: Page: 168FRA US/T Bu August 2001

ATA 76 ENGINE CONTROLS

THROTTLE CONTROL SYSTEM

GeneralThe throttle control system consist of :

− the throttle control lever− the throttle control artificial feel unit (Mecanical Box)− the thrust control unit− the electrical harness.

The design of the throttle control is based upon a fixed throttle concept :This means that the throttle control levers are not servo motorized.

Thrust Control UnitThe Thrust Control Unit contains two resolvers, each of which sends the thrustlever position to the Electronic Engine Control .The extraction current for theresolvers is provided by the EEC.

Autothrust Disconnect pushbutton.The autothrust instinctive disconnect pushbutton can be used to disengage theautothrust function.

THRUST LEVERS

GeneralThe thrust levers comprises :

− a thrust lever which incorporates stop devices and autothrust instinctive disconnect pushbutton switch− a graduated fixed sector− a reverse latching lever.

The thrust lever is linked to a mechanical rod. This rod drives the input lever ofthe throttle control artificial feel unit (Mechanical Box).

Reverse Thrust Latching LeverTo obtain reverse thrust settings, the revers thrust laching lever must be lifted.A mechanical cam design is provided to allow reverse thrust selection when-thrust lever is at fowward idle position.The thrust lever has 3 stops at the pedestal and 3 detents in the artificial feelunit: 0° STOP = FWD IDLE THRUST -20° STOP = FULL REVERSE THRUST 45° STOP = MAX .TAKE OFF THRUST DETENT � = (REVERSE) IDLE THRUST DETENT � = MAX.CLIMB (ALSO CRUISE SELECTION) DETENT � = MAX. CONTINOUS (FLEX TAKE OFF THRUST)

1

�

��

Page: 168

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ENGINE THRUST LEVER CONTROL

REVERSE THRUST

LATCHING LEVER

RESOLVER 1

RESOLVER 2

CHANNEL A

CHANNEL B

− FUEL METERING VALVE

AUTOTHRUSTDISCONNECT PB

THRUST LEVER

REVERSE THRUSTLATCHING LEVER

MECHANICAL BOX

THRUST CONTROL UNIT

CONTINUE

FMUEEC

Page: 169Figure 85 Engine Thrust Lever Control

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BUMP RATING PUSH BUTTON

This Push Buttons are optional equipment.In some cases the throttle control levers are provided with ”BUMP” rating pushbuttons,one per engine.This enables the EEC to be re-rated to provide addi-tional thrust capability for use during specific aircraft operations.

Bump Rating DescriptionThe takeoff bump ratings can be selected, regardless of the thrust lever angle,only in the EPR mode when the airplane is on the ground. The bump ratings, if available, are selected by a push button located on thethrust lever.Actuation of the switch will generate a digital signal to both EECs via the EIU.The maximum take-off rating will then be increased by the pre−programmeddelta EPR provided the airplane is on the ground.The bump ratings can be de−selected at anytime by actuating the bump ratingpush button as long as the airplane is on the ground and the thrust lever is notin the maximum takeoff (TO) detent.Inflight, the bump ratings are fully removed when the thrust lever is moved fromthe TO detent to, or below, the MCT detent.The bump rating is available inflight (EPR or rated N1 mode) under the following conditions. Bump rating initially selected on the ground. TO/GA thrust lever position set. Airplane is within the takeoff envelope.

The bump rating is a non−standard rating and is only available on certain designated operator missions.Use of the bump rating must be recorded.This information is for tracking bymaintenance personnel.

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Page: Page: 171FRA US/T Bu August 2001 Page: 171Figure 86 Bump Push Bottons

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ARTIFICIAL FEEL UNIT ( MECANICAL BOX )The Throttle control artificial feel unit is located below the cockpit center pedes-tal. this artificial feel unit is connected to engine 1(2) throttle control lever and tothe engine 1(2) throttle control unit by means of rods.The artificial feel unit is a friction system wich provides a load feedback to thethrottle control lever.This artificial feel unit comprises two symetrical casings, one left and one right.Each casing contains an identical and independent mechanism.Each mechanism is composed of:

− a friction brake assembly− a gear assembly− a lever assembly− a bellcrank assembly

Throttle lever travel is transmitted to the to the artificial feel unit and to thethrottle control unit.The linear movement of the throttle levers is transformed into a rotary move-ment at the belcrank wich turns about the friction brake assembly shaft. Thismovement rotates a toothed quadrant integral with the shaft.This toothed quadrant causes inverse rotation of a gear equipped with adisk which has four detent notches. Each notch corresponds to a throttlelever setting and is felt as a friction point at the throttle levers.

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MECHANICAL BOX(ES)MECHANICAL BOXES

ADJUSTMENT SCREW

RIGGING POINT

An adjustment screwis provided at thelower part of eachmechanical box to

DETENT FORCEADJUSTMENT

adjust the artificialfeel.

Page: 173Figure 87 Mechanical Boxes

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THROTTLE CONTROL UNITThe throttle control unit comprises : an input lever mechanical stops which limit the angular range 2 resolvers whose signals are dedicated to the EEC (one resolver per

channel of the EEC) 6 potentiometers fitted three by three. Their signals are used by the flight

control system a device which drives the resolver and the potentiometer a pin device for rigging the resolvers and potentiometers a safety device which leads the resolvers outside the normal operating

range in case of failure of the driving device two output electrical connectors.

The input lever drives two gear sectors assembled face to face. Each sectordrives itself a set of one resolver and three potentiometers.

Relation between TRA and TLA:The relationship between the throttle lever angle and throttle resolver angle(TRA) is linear and : 1 deg. TLA = 1.9 TRA.The accuracy of the throttle control unit (error between the input lever positionand the resolver angle) is 0.5 deg. TRA.The maximum discrepancy between the signals generated by the two resolversis 0.25 deg. TRA.The TLA resolver operates in two quadrants :the first quadrant serves for positive angles and the fourth quadrant for nega-tive angles.Each resolver is dedicated to one channel of the EEC and receives its electricalexcitation from the EEC.The EEC considers a throttle resolver angle value :

− less than −47.5 deg. TRA or− greater than 98.8 deg. TRA as resolver position signal failure.

The EEC incorporates a resolver fault accomodation logic. This logic allowsengine operation after a failure or a complete loss of the throttle resolver posi-tion signal.

Page: 174

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THRUST CONTROL UNIT(S)

RESOLVERRIGGING POINT

ELECTRICAL CONNECTORS

− 2 unitsEach unit consists of :− 2 resolvers− 6 potentiometers.

3 COUPLED POTENTIOMETERS

C

C

C

Page: 175Figure 88 Thrust Control Units

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RIGGINGThe throttle control levers must be at the idle stop position to perform the rigging procedure.

AIDS ALPHA CALL UP OF TRAUsing the Aids Alpha call up it is possible to check both TRA (Thrust ResolverAngle)

AIDS PARAM ALPHA CALL−UPENTER ALPHA CODE

− TRA EEC 1 :− TRA EEC 2 :

− ( )

− ( )

− ( )<RETURN PRINT>

0.00.1

Page: 176

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A

RIG PIN

THRUSTCONTROLUNIT

RIG PIN

MECHANICAL

BOX

Page: 177Figure 89 Thrust Control System Rigging

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AIDS ALPHA CALL UP OF TRAUsing the Aids Alpha call up it is possible to check both TRA (Thrust ResolverAngle)

Page: 178

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AIDS PARAM ALPHA CALL−UPENTER ALPHA CODE

− TRA EEC 1 :− TRA EEC 2 :

− ( )

− ( )

− ( )

− ( )

<RETURN PRINT>

0.00.1

Page: 179Figure 90 Alpha Call−up TRA

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ENGINE INDICATINGGENERAL

A319/A320/A321IAE V2530−A5

77−00


ATA 77 INDICATING

77−00 ENGINE INDICATING PRESENTATIONIndication GeneralThe engine is equipped with sensors that monitor

− temperature ,− pressure,− speed,− vibration− fuel flow

lt also has switches that provide indication for− oil, fuel clogging.− thrust reverser hydraulic pressure.− position (SAV, T/R, Overspeed governor, etc....

Depending on the data transmitted, messages are generated on the followingdevices :

− Upper ECAM : Engine Warning Display (EWD).− Lower ECAM :Systems Display (SD).− Master caution, or warning.− Audible chimes and oral warning.

These messages are used to run the engine under normal conditions through-out the operating range, or to provide warning messages to the crew and main-tenance personnel. The master caution and warning are located in front of thepilot on the glance panel.

Primary Engine DisplayThe primary engine parameters listed below are permanently displayed on theEngine and Warning display ( E/WD ): Engine Pressure Ratio ( EPR ) Exhaust Gas Temperature ( EGT ) N1 ( low rotor speed ) N2 ( high rotor speed ) FF ( fuel flow )

After 5 min of the power up test the indication is displayed in amber and figuresare crossed ( XX ). Normal indication can be achieved by using the FADECGRD power switches, one for each engine at the maintenace panel or by the

MODE selector switch on on the Engine panel at the pedestal in CRANK orIGN / START position for both engine.If a failure occurs on any indication displayed, the indication is replaced by am-ber crosses, the analog indicator and the marks on the circle disappear, thecircle becomes amber.Only in case of certain system faults and flight phases a warning message ap-pears on the Engine Warning Display.

Secondary Engine DisplayThe lower display shows the secondary engine parameters listed below. Theengine page is available for display by command, manually or automaticallyduring engine start or in case of system fault: Total FUEL USED

For further info see ATA 73 OIL quantity

For further info see ATA 79 OIL pressure

For further info see ATA 79 OIL temperature

For further info see ATA 79 Starter valve positions, the starter duct pressure and during eng start up,

that operating Ignition system ( ONLY ON ENGINE START PAGE ) In case of high nacelle temperature a indication is provided below the en-

gine oil temp. indication. Engine Vibration − of N1 and N2 As warnings by system problems only:

− OIL FILTER COLG− Fuel FILTER CLOG

Some engine parameters also displayed on the CRUISE page

Page: 180

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77−00


FF KG / H

FOB: 19.125

� � � � � � � � � ��

35 35

IGNA BONLY ON ENGINE

START PAGE

Page: 181Figure 91 Engine ECAM Indications

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Page: 182FRA US/T JaG MAR 2006

ATA 77−10 POWER INDICATINGEPR INDICATION

EPR − Engine Pressure RatioThe Engine Pressure Ratio indicating system consists of one combined P2 / T2sensor and eight ports located in each of the three LPT exhaust case struts,P4.9.The pressure from this sensors are routed to the EEC pressure transducer.TheEEC converts the signal to a digital format and proccess the pressure to formactual EPR ( P 4.9 / P 2 ) and transmits the EPR value to the ECAM. Each ofthe two channels performs this operation independently.

1 Actual EPR

Actual EPR is green.

2 Cyan EPR command arc ( transient )

from current EPR pointer to EPR command value. is only displayed with A / THR engaged.

3 EPR TLA ( white circle )

Predicted EPR corresponding to the thrust lever position.

4 EPR max ( thicker amber mark )

It is the limit value of EPR corresponding to the full forward thrust lever position.

5 REV indication

Appears in amber when one reverser is unstowed or unlocked or inadvertenly deployed. ( In flight, the indication first flashes for 9 sec. and then remains steady. It changes to green when the reverser is fully deployed .

6 Thrust limit mode, EPR rating limit TO GA, FLX, MCT, CL, MREV selected mode is displayed in green, the associated EPR rating is displayed in blue. In MREV no EPR value is dis played.

Thrust limit mode is displayed in digital form, it indicates the mode which theEPR limit value will be computed.

− In flight ( or on ground with ENG stopped ): The selected mode corresponds to the detent of the most advanced

thrust lever position Rating limit is computed by the EEC receiving the highest actual EPR

value ( exept on ground with ENG stopped where it is computed bythe EEC receiving the most advanced thrust lever position ).

NOTE: WHEN A THRUST LEVER IS SET BETWEEN TWO POSITIONSTHE EEC SELECTS THE RATING LIMIT CORRESPONDING TOTHE HIGHEST MODE.

NOTE: WHEN IDLE IS SELECTED THE EEC SELECTS CL

NOTE: WHEN M REV IS SELECTED, THE EPR RATING LIMIT VALUE ISREPLACED BY AMBER CROSSES ( M REV MODE IS LIMITED BYN1)

− On ground ( with engines running ) With engines running, on ground, whatever the lever position is,this

limit corresponds to: TO GA thrust limit. With engine running, on ground, if FLX mode is selected, FLX EPR is

displayed whatever the thrust lever position between IDLE and FLX /MCT.

If FLX mode is selected, the flexible take off temperature in � C, selectedthrough the FMS MCDU’ s, is displayed. For FLX mode indication the ADIRU‘smust be switched on.The temperature value is displayed in green and the � C is displayed in blue.If a failure occurs on any indication displayed, the analog indication is replacedby amber crosses, the analog indicator and the marks on the circle disapear,the circle becomes amber.

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Page: 183FRA US/T JaG MAR 2006

TOGA 1. 520

FLX 1.503 35� C

MCT

EPR

CL

MREV

OR

OR

OR

OR

REV

1

32 2

4 4

5

66

3

Page: 183Figure 92 EPR Indication − Upper ECAM Display Unit

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EPR SYSTEM COMPONENTS

P2 / T2 SENSORThe P2 / T2 sensor is located near the 12 o’clock position of the inlet cowl. Itmeasures total pressure and temperature in the inlet air stream of the engineforward of the engine front flange. The dual output total temperature measure-ment is accomplished by two resistance−sensing elements housed in theP2/T2 sensor body. Each channel of the Electronic Engine Control ( EEC )monitors one of these resistance elements and converts the resistance mea-surement to a temperature equivalent.The total air pressure is carried via pressure tubing to the pressure sensor lo-cated in channel A of the EEC.The P2 / T2 sensor has an anti−icing function accomplished by a single heat-ing element internally bonded to the sensor. The heater is a hermeticallysealed, coaxial resistance element brazed internally to the sensor casting. Air-craft power, which is used for the heater, is switched on and off by the EECdepending on TAT ( < 7,2 ° C heater ” ON ” ), via the relay box.

NOTE: IN CASE OF LOSS OF P2 / T2 HEATING , AN AUTOMATIC RE-VERSION FROM EPR MODE TO UNRATED N1 MODE OCCURS.

P4.9 SENSORS

The P4.9 sensor and manifold has three probes which measure the total pres-sure of the exhaust gas stream.Struts 4, 7 and 10 contain the pressure sensing ports. Each sensing point con-tains eight radial pressure sensing ports which are combined to yield an aver-age pressure. The resulting average radial pressure value from each strut isthen plumbed into a manifold which provides an overall turbine exhaust pres-sure average ( P4.9 ). A tube from this manifold is connected to the ElectronicEngine Control ( EEC channel A ).A pressure transducer located within the EEC converts the average pressure atstation 4.9 into a useable electronic signal ( proportional to pressure ) that canbe processed and used by the EEC to control the engine.

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P2/T2 SENSOR

P 4.9 SENSOR

PRESSURE CONTROLMANIFOLD

PRESSURE CONTROLMANIFOLD

ADAPTOR

P4.9 PRESSURERAKE TUBE

P2/T2 SENSOR

A

B

B

Page: 185Figure 93 P2 / T2 and P4.9 Sensor

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P2 / T2 HEATERAircraft Power ,which is used for the heater , is switched on and off by theEEC, via the relay box.The heater and the heating Circuit can be tested using the FADEC CFDS Testmenu.

NOTE: THE RELAY BOX ALSO CONTAINS THE 115V IGNITION RELAYS.

FAIL SAFE POSITION: ”PROBE HEATER OFF”

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RELAY BOX

CH BCONNECTOR

CONNECTOR

CH A

P2/T2 HEATINGCONNECTOR

RELAY BOX

INPUT FORIGNITION RELAYS

11DA2 C/BANTI ICE / PROBESP2/T2 ENG 2122VU212

204XP-C115VACBUS 224-58-06

1WDENG/APU FIRE PNL20VU210 26-12

4100KSRELAY BOX446STA450 73-25

4014KSSENSOR P2/T2

444STA390 73-25

Page: 187Figure 94 P2/T2 Heater Schematic

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FADEC P2/T2 HEATER TEST

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Page: 189FRA US/T Bu August 2001 Page: 189Figure 95 P2/T2 Heater Test

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ATA 77−20 TEMPERATURE

EGT INDICATION

EGT Indicator

1 Actual EGT

Normally displayed in green. Pulses amber up to MCT when EGT � 610 � C. Pulses red when EGT �650 � C.

NOTE: EGT INDEX PULSING AMBER MUST BE DISREGARDED WHENUSING TO OR FLX THRUST.

2 Max EGT

Thicker amber mark is set at � 610 � C, it is the max EGT value up to MCT thrust. It is not displayed during: −Engine start up, instead a amber mark is placed at 635 � C −Take Off sequence.

3 Max permissible EGT

Goes up to 650 �C. A red band begins at the point of over temperature and a red cross line appears at the max value achieved.

4 Red cross line

is set at the max EGT over temperature achieved during the last leg. The red cross line will disappear through corresponding DMC’ s − MCDU action or by the next T/ O.

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1

2

34

Page: 191Figure 96 EGT Indication

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EGT PROBES

The messurement channel for the exhaust gas temperature consist of: Four probe assemblies, each comprizing 2 thermocouples.

− four thermocouples ( one from each probe assembly ) are used toform an averaged signal send to the channel ” A ” of the EEC.

− the remaining four thermocouples ( one from each probeassembly ) are used to form an averaged signal, send to channel ” B ” of the EEC.

The EEC uses the Exhaust Gas Temperature in the engine start control logicand also transmits the EGT signal to the ECAM .

The EGT probes are located at engine station 4.95 ( LPT exhaust case strut ),at 9.5, 7.5, 4.5 and 2 O ’Clock.The thermocouples are connected, in parallel, to the junction box for eachchannel, from where two indepent signals are send to the EEC. Each signal isan average of the four probes.

9.5

7.5 4.5

2.0

JUNCTION BOX

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1 CHROMEL STUD

EGT JUNCTION BOX

JUNCTION BOX

Page: 193Figure 97 EGT System

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ATA 77−10 POWER

N1 AND N2 INDICATION

N1 IndicationThe low pressure rotor speed signal is used in the EEC for engine control com-putation and for ECAM visual display.

1 Actual N1

Displayed normaly in green. Pulses red if N1 exceeds 100%. Pulses amber when N1 exeeds the N1 rating limit, in N1 MODE.

2 Max permissible N1

is 100 %. At 100 % a red band begins. If the RPM exeeds 100 % index and numeric value pulses red.

3 Red cross line

is set at the max N1 over speed value achieved during the last leg.

4 White circle

N1 command corresponding to the thrust lever ( angle ) position ( predict N1 ) appears when in rated N1 mode. N1 rated MODE can activated automaticly or by switching the N1 MODE switch at the overhead panel ( close to the ENG MAN START switches ). Both engine must be in the same MODE, rated or unrated. Not displayed in unrated N1 MODE. Auto thrust is not active in rated N1 mode .General: A failure title will be displayed on E / WD in the MEMO display.

5 CHECK

appears for EPR, EGT, N1 , N2 and FF, if the displayed value compared by the DMC’ s with the actual value from the EEC differs and the last digit from the value shown will be XX ed.

6

6 N1 MODE switches

ON: −Thrust control reverts from EPR mode to N1 rated mode. Following an automatic reversion to N1, rated or unrated mode, pressing the P/B switch to confirm the mode. ON, it illuminates blueOFF: − If available, EPR mode is selected

N2 IndicationThe signal fore the HP rotor speed is originated from the dedicated alternatorto the EEC for use in engine control computation and to the ECAM for visualdisplay on ECAM. A separate signal goes to the engine vibration monitoringunit ( EVMU ) for use in processing engine vibration data.

7 Actual N2

Digital indication normally green. It is overbrightness and grey boxed during engine start sequence up to 43 % ( starter cut out ). Turns red if N2 exceeds 100 % and a red ” X ” appears. The red ” X ” will disappear through corresponding DMC’s − MCDU action or by the next T/O.General: A failure title will be displayed on E / WD on the MEMO display.

If a failure occurs on any indication displayed, the analog indication is replacedby amber crosses, the analog indicator and the marks on the circle disapear,the circle becomes amber.

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1 2

3

4

ENG 1 EPR MODE FAULT

MCTN 1 MODE

95.8XX XXEPR

CHECK

5

4

4

7 X

7 ENG 1 N 2 OVER LIMIT

Page: 195Figure 98 N1 and N2 Speed Indication

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INDICATINGGENERAL

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Page: 196FRA US/T kh August 2001

ATA 31 INDICATINGMAX POINTER RESET ( N1, N2 & EGT )

Monitoring of the relevant display of the engine parametersN1 , N2, EGT, and FF indications of both engines are monitored internally andexternally .The DMC compares the N1 signal received from the EEC 1 with thefeedback signal which reflects the displayed position of the N1 needle −In order to grant dissimilarity with the engine 2 monitoring process the DMCcompares the N1 signal from the EEC 2 with the feedback signal representingthe N1 digital value.The same applies to the EGT parameters indications, but with the displayedposition of the engine 2 EGT needle and the engine 1 EGT digital feedbackvalue.As for the N2 and FF parameters, the DMC compares the direct signal from theEEC with the displayed digital value.In case of detected discrepancy, a CHECK amber message is displayed justbelow the relevant parameter indication .In addition the FWC,s perform an external monitoring between the feedbacksignals (that correspond to the displayed values and the signets that are di-rectly received by the FWC’s from the EEC‘sShould a descrepancy occur, for one or more parameters, a CHECK ambermessage is displayed under the relevant indicationThe FWC’s generate a caution

− single chime− master caution Light− message on the upper ECAM DU : ENG 1 (2) N1(N2/EGT/FF) DISCREP- ANCY

Max pointer Reset ( N1, N2 & EGT )The Max pointers for N1, N2 and EGT can be reset using the CFDS menuINSTRUMENTS. The menu for the EIS 1,2,3,( DMC 1,2,3 ) must be selectet.The memory cells which store the possible exceedance are reset either bypressing the GENERAL RESET line key or automatically at the next take off.

Read−out/Reset of the Engine Red Line ExceedancesThe DMC connected to the upper ECAM DU monitors primary parameter indications of both engines.Should an exceedance occur, the DMC memorizes in its BITE memory themaximum value reached during the Last Flight LegThe values of the N1, N2, EGT red lines and transitory overlimit values arestored in 2 independent tables, one per engine.Read out of this engine parameter exceedance can be performed via the DMCMCDU menu.With the function engines the parameters can be selected eitherfor engine 1 or 2.

NOTE: A RESET OF THE RED LINE LIMITS HAVE TO BE PERFORMEDON ALL 3 DMCS.

N1 RED LINE ExceedanceThe N1 red line is represented by an arc shaped red ribbon situated at the endof the scale.If the N1 actual value exceeds the N1 red line (even for a short period of time),a small red line appears across the N1 scale and then stays at the maximumvalue which has been reached.This indicates a N1 exceedance condition.Should this condition occur, thesmall red line disappears only after a new take−off or after a maintenance ac-tion through the MCDU DMC reset.

N2 RED LINE ExceedanceThe N2 indications are displayed in digital form only. 100% N2 correspond to14460 RPM.Should N2 actual exceeds the N2 red line value,a red cross ap-pears next to the digital indication. This red cross disappears only after a newtake off or a DMC reset.

EGT RED LINE ExceedanceThe EGT indications are provided in the same form as for the N1 indications.The same applies to changes in color and EGT exceeding indications.However it has to be noticed that the amber linie (EGT MAX) is vari-able.635 deg. C at engine start and 610 deg. C afterwards.Red line Limit is 650deg.C.

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Page: 197FRA US/T kh August 2001 Page: 197Figure 99 Max Pointer Reset

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ATA 77-10 POWER

N1 INDICATION

The fan speed ( N1 ) indication system has four sensors:

Two of them are used to provide EEC channels ” A ” and ” B ” with N1 ro−tational speed signal.

One sensor acts as a spare fore either EEC channel ( it can be activated bychangeover connectors at the junction box ).This sensor cannot be used in place of the N1 sensor dedicated to the En−gine Vibration Monitoring Unit with N1 analog signals ( trim balance sen−sor ), see below.

One sensor provides the Engine Vibration Monitoring Unit with N1 analogsignals ( trim balance sensor ).

The N1 electrical harness tube goes through the inner strut of the no. 3 strutof the intermediate structure and to the terminal block.The electrical leads from each sensor goes through the N1 tube and is con-nected to the terminal block.

For the fan speed sensors, one turn on the LP shaft causes 60 teeth on thephonic wheel to pass its sensor.For the trimbalance sensor, one slot in the phonic wheel passes the sensorone time for one turn.

The EEC speed sensors have two pole pieces compared to the trimbalancesensor who has only one pole piece.

INTERCHANGE OF N1 SPEED SENSORSTask 77−11−00−860−010 If the fan speed sensor No. 1 is unserviceable, disconnect the harness

leads No. 1 and No. 2 from their terminals No 1 and No 2. Reconnect the harness lead No 1 to the terminal No. 3 and the harness lead No. 2 to the terminal No. 4 of the spare speed sensor. If the fan speed sensor No 3 is unserviceable, disconnect the harness leads

No. 5 and No. 6 from their terminals No. 5 and No. 6 and reconnect the har-ness leads to the spare speed sensor as described above.

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THREE FAN SPEED SENSORS

TWO POL PIECES

TERMINALNO. 4 (SPARE)

TERMINALBLOCK

ONE TRIM BALANCE SENSOR

ONE POL PIECE

Page: 199Figure 100 Fan Speed & Trim Balance Sensor,N1 Terminal Block

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DEDICATED ALTERNATOR (PMA)

The alternator function are: the primary power source for the Electronic Engine Control (EEC) N2 signal source for the EEC and Engine Vibration Monitoring Unit

(EVMU) and the cockpit

DescriptionThe unit is designed for maximum reliability by the elimination of splines, bear-ings or similar parts which can deteriorate or fail.The rotor is mounted directly on the gearbox output shaft and the stator isbolted to the gearbox housing.

The alternator provides two identical and independent power outputs, onefor each channel of the EEC. It comprises two stators (one power and one speed) and a rotor. Is driven from the main accessory gearbox Consists of a magnetic rotor running in a stator. the stator has four inde-

pending windings, two of which provide three phase frequency AC electricpower to respectively channel ” A ” and ” B ”.The third winding provides a single phase AC analog signal propotional toN2 for the Engine Vibration Monitoring System. The forth winging provides a dedicated N2 signal to Channel ”A ” of theEEC.

The N2 windings gives an analog signal through the cockpit for ECAMindication.

The stator and rotor are sealed from the gearbox by a shaft seal. If a shaft sealfailure occurs and the alternator fills with engine oil, the alternator will continueto function normally.To maintain the temperature of the dedicated alternator at an acceptable levelthe alternator incorporate an integral cooling air manifold using fan air.

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P12,5 AIRA

A

Page: 201Figure 101 Engine Dedicated Alternator

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ENGINE INDICATINGANALYZERS

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VIBRATION INDICATIONAn engine vibration monitoring unit monitors the N1 and N2 levels of both engines.

GeneralThe engine vibration measurement system comprises :

− one transducer on each engine with 2 piezoelectric accelerometers .− an Engine Vibration Monitoring Unit− two vibration indications N1 and N2.

The engine vibration system provides the following functions :− vibration indication due to rotor unbalance via N1 and N2 slaved tracking filters− excess vibration (above advisory level of 5 units )− fan balancing (phase and displacement)− shaft speed (N1 and N2)− storage of balancing data− initial values acquisition on request (option )− BITE and MCDU communication− accelerometer selection− frequency analysis when the printer is available.

NOTE: ONLY ONE ACCELEROMETER IS USED AT A TIME (A OR B).THE SAME ACCELEROMETER IS NOT USED FOR TWOSUCCESSIVE FLIGHTS. THE CHANGEOVER OCCURS AT POW-ER−UP OR ON SPECIAL REQUEST (MCDU) ON THE GROUND.

InterfacesThe EVMU interfaces with the ECAM and the CFDSCFDS interfaces: Maintenance fault messages.The N1 and N2 vibrations of the left and right engines are displayed on the en-gine and cruise pages.

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0.8

1.2

0.8

1.2

CFDIU

SDAC2

SDAC1

VIBRATION indications:

VIB N1

VIB N20.8 0.9

THE VIBRATION INDICATIONSOF THE LP AND HP ROTORS AREDISPLAYED IN GREEN.

PULSINGADVISORYABOVE 5

PULSINGADVISORYABOVE 5 1.2 1.3

80 80

140 160

VIB SENSOR A

VIB SENSOR B

DED: GEN

Powersupply115V AC

DMU AIDS

Page: 203Figure 102 Vibration Indication

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ENGINE VIBRATION MONITORING UNIT (EVMU)

DescriptionThe signal conditioner is composed of:

− 2 channel modules− 1 balancing module− 1 data processing module− 1 power supply module.

These modules are removable parts from the signal conditioner and are repairable subassemblies.

Channel modulesEach channel module processes the signals from the two engineaccelerometers and from the two speed signals N1 and N2 : this enablesthe extraction from the overall vibration signal of a component due torotor first order unbalance.The N1 and N2 signals are used to:

− drive the tracking filters, and− slave their center frequencies at the shaft rotational speed.

The accelerometer signals pass through these tracking filters whichextract the N1 and N2 related fundamental vibration. The accelerationsignal is then integrated in order to express the vibration in velocityterms.The EVMU receives analog signals from :

− the 2 engine accelerometers (1 per engine)− and the N1 and N2 speed sensors of each engine.

It also receives digital input from CFDS through ARINC 429 data bus.The EVMU sends signals through the digital ARINC 429 data bus to :

− SDAC1 and 2 for cockpit indication− the CFDIU− the DMU− and printer (if installed) for maintenance purposes.

Power supply moduleThe power supply module receives the 115VAC/400Hz power. It providesthe other modules with the necessary voltages.

Power SupplyThe EVMU is supplied with 115V/400Hz by the busbar 101XPA, through thecircuit breaker 1EV.

Built in test equipment (BITE) maintenance and fault informationThe equipment contains a BITE system to detect internal and external failure.During the execution of the cyclic BITE sequence, the following partsof the EVMU are checked:

− the non−volatile memory− the timers− the analog−to−digital converter− the ARINC 429 transmitter and receivers− the tacho generators.

During the power−up sequence of the BITE, the following parts of theEVMU system are checked:

− N1 and N2 NB velocity− unbalance data− N1 and N2 tacho frequencies− accelerometer signals.

Any detected failure is stored in the non−volatile memory with GMT, the dateand other reference parameters.

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Page: 205FRA US/T Bu August 2001 Page: 205Figure 103 EVMU Schematic

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Page: 206FRA US/T KoA May 04

COMPONENTSThe vibration transducer including two indipendent channels is installed on thefan case at the top left side of the engine.The EVMU is located in the Avionics compartment 86VU.

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Page: 207FRA US/T KoA May 04

VIBRATION TRANSDUCER

FAN CASE

ELECTRICAL HARNESS

VIBRATION SENSOR

Page: 207Figure 104 Vibration Sensors

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Engine IndicatingAnalyzers

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Page: Page: 208FRA US/T bu July 01

EVMU OPERATION (CFDS)

The maintenance mode (CFDS mode) of operation, through the MCDU menu,allows maintenance staff to obtain, or print the followingFirst page :

− Last leg report : lists the LRU’s detected faulty during the last leg.− Previous leg report : lists the LRU’s detected faulty during the legs (max 62) previous to the last leg.− LRU identification : provides the unit part number and manufacturer name.− Class 3 failures : lists the LRU’s detected faulty during a ground test. Only the last 3 failures detected are displayed.− Test: enables a complete check of the EVM system. lf no failure has been detected, the message TEST OK is displayed. lf any failure has been detected the failed LRU is displayed.

Second page:− Accelerometer reconfiguration : allows selection of the accelerometer (fan no. 1 bearing or TRF) to be used for the next flights. The EVMU also indicates which accelerometer is in operation.− Engine unbalance : allows selection of 5 different speeds per engine (from 50% to 100% N1 RPM) at which unbalance data is stored. lt also enables previously ac- quired unbalance data to be read, and performs balancing calcula- tions for both engines using both accelerometers.− Frequency analysis : enables a frequency analysis of the acceleration signal to be per formed. The results are sent to the printer.

− Shop maintenance : operational only when the EVMU is in the shop for maintenance. lt allows troubleshooting of the LRU itself and corresponds to failures at LRU level.

Page: 208

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77−30

Page: Page: 209FRA US/T bu July 01 Page: 209Figure 105 EVMU CFDS Pages

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77-30


CFDS SYSTEM REPORT / TEST

The Centralized Fault Data System (CFDS) enables access to the system.The first menu sent to the MCDU is the main menu. The various functions aredetailed here after.

Last leg reportThe EVMU sends the list of the LRUs which have been detected faulty duringthe last leg.

Previous leg reportThe EVMU sends the list of the LRUs which have been detected faulty duringthe legs (maximum 64) previous to the last leg. The faults detected are thesame as for the last leg report.

LRU identificationThe EVMU sends the EVM unit part number

TestThe test item allows initiation of a complete check of the EVM system.If no failure has been detected, the message ”TEST OK” is displayed.If any failure has been detected the failed LRU is displayed.

Page: 210

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TEST TEST

TEST IN PROGRESS

TEST

SELF-TEST O.K.

OR

77-32-16ENG1 ACCLRM 4004EV (A)

Page: 211Figure 106 CFDS System Report / Test EVMU

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CFDS SYSTEM REPORT /TEST

ENGINE UNBALANCE MENUThis menu permits for both engine, to command unbalance data storage during next flight and the read out of the stored data. It also permits to effectuate balancing for a selected engine with both accelerometers.

Measurement of the unbalance dataThe EVMU measures the position and the amplitude of the rotor unbalance ofeach engine. It provides this information, when available,to the output bus.

Storage of unbalance dataIf requested, the system can store the balancing data during the cruise phasewhen stabilized conditions are reached (the actual N1speed does not fluctuatemore than plus or minus 2% during at least 30s). For every stored measure-ment the stabilized conditions shall be met once more again.

NOTE: THIS TEST CAN BE DONE DURING AN ENGINE RUN-UP IN OR-DER TO OBTAIN VIBRATION MEASUREMENT FOR DIFFERENTN1 SPEEDS. REFER TO AMM ATA 77-32-34.TO GET ACCESS AGAIN TO THE SYSTEM REPORT / TESTMENU ENG, REFER TO AMM 31-32-00.

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EVMUBALANCING LEFT

< ACC.A START ACC.B >

20 / 59 N 1 / N 2 % 20 / 59

PHASE DEG359 3590 / 359 0 / 359

0.1 DISPL MILS 0.00.1 / 0.1 0.0 / 0.1

< ACC.A * ACC.B >STOP

EVMUBALANCING LEFT

< ACC.A START ACC.B >

00 / 00 N 1 / N 2 % 00 / 00

PHASE DEG0 00 / 0 0 / 0

0.0 DISPL MILS 0.00.0 / 0.0 0.0 / 0.0

< ACC.A * ACC.B >STOP

EVMU

< RETURN PRINT >

ENGINE UNBALANCE

< LEFT

< LEFT

< LEFT

RIGHT >

RIGHT >

RIGHT >

READ

BALANCING

EVMU

<ACCELEROMETER RECONFIGURATION

< ENGINE UNBALANCE

< FREQUENCY ANALYSIS

CLEAR

NOTE: THE N1 SPEED CAN BE INDI-CATED IN % OR RPMDEPENDING ON EVMU SOFT-WARE.

Page: 213Figure 107 Unbalance Data

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CFDS SYSTEM REPORT /TEST

ENGINE UNBALANCE MENUThe EVMU acuired unbalance data can be cleared with the clear menu.

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Page: 215FRA US/T Bu August 2001 Page: 215Figure 108 Unbalance Data

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CFDS ACCELEROMETER RECONFIG.This menu allows selection of the accelerometer A or B or the auto switchmode alternate to be used for the next flights.The EVMU indicates which accelerometer is in operation.

Page: 216

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Page: 217FRA US/T Bu August 2001 Page: 217Figure 109 Accelerometer Reconfiguration

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AIRCRAFT INTEGRATED DATA SYSTEM

AIDS reportsVibration data is provided to the Aircraft Integrated Data System (AIDS), whichis used to monitor aircraft and engine parameters.lt allows maintenance staff to perform engine parameter trend monitoring andtroubleshooting.The vibration information is printed on various reports, which are:

− Engine cruise report.− Cruise performance report.− Engine take−off report.− Engine on−request report.− Engine mechanical advisory report.− Engine run−up report.

Page: 218

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Page: Page: 219FRA US/T bu July 01 Page: 219Figure 110 AIDS

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ENGINE FUEL AND CONROLCONTROLLING

A319/A320/A321IAE V2530 −A5

73−20



73−20 FADEC

FADEC LRU’S

Electronic Engine Control (EEC )

Fuel Metering Unit ( FMU )

Sensors

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IDGTOIL

GEN

/PB

Page: 221Figure 111 FADEC Architecture

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FADEC LRU‘S

Electronic Engine Control ( EEC )

Data Entry PlugThe Data Entry Plug ( DEP ) provides discrete inputs to the EEC.Located tothe Junction 6 of the EEC it provides unique engine data to channel A and B. The data transmitted by the DEP is: EPR Modifier (Used for power setting ) Engine Rating Engine Serial No.

NOTE: IF THE DATA INPUTS OF THE DATA ENTRY PLUG J6 ARE LOST,THEN AN AUTOMATIC REVISION FROM EPR MODE TO UN-RATED N1 MODE OCCURS.

DATA ENTRY PLUG MODIFICATION

DescriptionThe DEP links the coded data inputs through the EEC by the use of shortingjumper leads which are used to select the plug pins in a unique combination.During a life of an engine , it may be necessary to change the DEP configura-tion , either during incorporation of Service Bulletins or after engine overhaul ,when the EPR modifier code may need to be changed.This is accomplshed bychanging the configuration of the jumper leads in accordence with the relevantinstructions.During removal/replacement of the DEP it is necessary to use an EEC HarnessWrench as it is imperative that the connectors are tight. On fitment of the DEPto the EEC align the main key of the connector with the EEC and hand tightenthe connector.Then using the EEC Harness Wrench torque tighten the DEPconnector to 32 Ibf in.

NOTE: THE PART NUMBER IS WRITTEN ON THE DEP . THE PARTNUMBER CAN ALSO BE FOUND ON THE ENGINEDATA PLATE,WHICH IS LOCATED AT THE LEFT HAND SIDE OFTHE FAN CASE.

EEC DEP TESTERAfter modifing the DEP a electrical wiring test on the data entry plug assembly-must be performed with the tester below,to make sure the pins and jumpers areproberly installed.

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EEC

DATA ENTRYPLUG

CHANNEL B HOUSINGCHANNEL A HOUSING

VIBRATION ISOLATORMOUNTS HANDLE

COOLING AIR PORTS

PRESSURE PORTS

*MARKING AERA

Page: 223Figure 112 EEC/ Data Entry Plug

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ELECTRONIC ENGINE CONTROL (EEC)

Harness (Electrical) and Pressure ConnectionsTwo identical, but separate electrical harnesses provide the input/output circuitsbetween the E.E.C. and the relevant sensor/control actuator, and the aircraftinterface.The harness connectors are ’keyed’ to prevent misconnection.

NOTE: SINGLE PRESSURE SIGNALS ARE DIRECTED TO PRESSURETRANSDUCERS − LOCATED WITHIN THE E.E.C. − THE PRESSURE TRANSDUCERS THEN SUPPLY DIGITAL ELECTRONIC SIGNALS TO CHANNELS A AND B.

The following pressures are sensed:. Pamb - ambient air pressure ( fan case sensor ). Pb - burner pressure (air pressure) P3/T3 probe. P2 - pressure ( P2/T2 fan itlet probe ). P2.5 - booster stage outlet pressure. P5 (P4.9) - L.P. Turbine exhaust pressure ( P5 (P4.9) rake ). P12.5 - fan outlet pressure ( fan rake )

Electrical Connections

Front FaceHarness Connector Plug Identification

J1 E.B.U. 4000 KSAJ2 Engine D202PJ3 Engine D203PJ4 Engine D204PJ11 Engine D211P

Rear FaceJ5 Engine D205PJ6 Data Entry PlugJ7 E.B.U. 4000 KSBJ8 Engine D208PJ9 Engine D209PJ10 Engine D210P

Page: 224

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FRONT FACE REAR FACE

BOTTOM FACE

Page: 225Figure 113 Electronic Engine Control ( EEC )

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ENGINE AND FUEL CONTROLCONTROLLING

A319/A3207A321IAE V2530 A5

73−20


FADEC POWER SUPPLY

EIU Power supplyThe EIU is powered from the aircraft electrical power, no switching has to bedone.

Electronic Engine Control (EEC) Power SupplyThe EEC is supplied from the aircraft electrical power when engine is shut-down, then from the EEC generator when the engine is running.

− aircraft electrical power when N2 <10%.− EEC generator power when N2 >10%.

Powering N2 <10%Each channel is independently supplied by the aircraft 28 volts through the En-gine Interface Unit.A/C 28 VDC permits :

− automatic ground check of FADEC before engine running− engine starting− powering the EEC while engine reaches 10% N2.

NOTE: THE EIU TAKES POWER FROM THE SAME BUS BAR AS THEEEC.

Powering N2 >10%As soon as engine is running above 10% N2, the EEC generator can supplydirectly the EEC.The EEC generator supplies each channel with three−phase AC. Two TRU’s inthe EEC provides 28VDC to each EEC channel.

Auto DepoweringThe FADEC is automatically depowered on ground, through the EIU after en-gine shutdown.EEC automatic depowering on ground :

− after 5 mn of A/C power up.− after 5 mn of engine shutdown

NOTE: AN ACTION ON THE ENG FIRE P/B PROVIDES EEC POWER CUTOFF.

FADEC Ground Power PanelFor maintenance purposes and MCDU engine tests, the FADEC Ground PowerPanel permits FADEC power supply to be restored on ground with engine shutdown.When the corresponding ENG FADEC GND POWER P/B is pressed ”ON” theEEC is powered again .

NOTE: ALSO THE FADEC IS REPOWERED AS SOON AS THE ENGINEMODE SELECTOR OR THE MASTER LEVER IS SELECTED .

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401 PP (DC ESS BUS)FOR ENGINE 1 & 2

202 PP (DC BUS 2 )FOR ENGINE 2

301 PP (BAT BUS)FOR ENGINE 1

NOTE: * supplied for 5 min

EEC

EEC

DEDICATEDGEN 28V

A

B

TRU/

28VTRU/

NORM

Page: 227Figure 114 FADEC Power Supply

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A319/A3207A321IAE V2530 A5

73−20


49VU

2450000HMQ0


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121VU

122VUANTI ICE

2450000UMR0

2450000VAQ0


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ENGINE FUEL AND CONTROLFADEC SENSORS

A319/A320/A321IAE V2530 A5

73−20


ATA 73-22 FADEC SENSORS

FADEC LRU‘S SENSORS

Engine Sensors

T4.9 (EGT) Sensor

(Ref. 77−20−00)

N1 Sensor

(Ref. 77−10−00)

N2 Sensor

(Ref. 77−10−00)

Engine Oil Temperature Sensor

(Ref. 79−30−00)

P2/T2 Sensor

(Ref. 77−00)

P3/T3 Sensor

P4.9 ( P5)

Page: 230

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P2/T2

T4.9

T4.9

P4.9 (P5)

P3 / T3

N2

P12.5

T2.5

P4.9 (P5)

T4.9

T4.9P0 (Pamb)

EEC

P2.5

N1

P4.9 (P5)

Page: 231Figure 117 FADEC Sensors

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FADEC LRU‘S SENSORS

P3/T3 SENSORThe P3/T3 sensor monitors the pressure and temperature at the exit of the HPcompressor.The combined sensor houses two thermocouples and one pressure inlet port.Each thermocouple provides an independant electrical signal, proportional totemperature, to one channel of the Electronic Engine Control (EEC).

PURPOSE:The purpose of the P3/T3 sensor is to provide performance data to the EEC forstarting and during transient and steady state operation of the engine.

P3/T3

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PRESSUREPORT

CHROMEL

ALUMEL

Page: 233Figure 118 P3/T3 Sensor

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P12.5 SENSORThe P12.5 sensor is a pressure tapping at the top of the fan case. It monitorsthe pressure behind the fan stator. This pressure is used for trend monitoring.The pressure tapping is also used for the cooling air supply of the dedicatedalternator(see Fig.114).

P2.5 / T2.5 SENSORSThese two sensors are located in the intermediate case. They are monitoringthe pressure and temperature between the two compressors. T2.5 is used forsystem scheduling, P2.5 is used for trend monitoring.

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P12.5 OFFTAKE

Page: 235Figure 119 P2.5 / T2.5 Sensors

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FADEC DESCRIPTION

GeneralThe Full Authority Digital Engine Control system consists of an Electronic En-gine Control plus a Fuel Metering Unit,sensors and peripheral components.

Electronic Engine ControlThe EEC consists of two channels (A and B) with crosstalk.Each channel cancontrol the various components of the engine systems.They are permanently operational. one channel is in command while the otheris in standby . In case of failure of the operational channel, the system auto-matically switches to the other one.

NOTE: THE CHANNEL SELECTION STRATEGY IS BASED ON CHANNELHEALTH CRITERIA .THE COMMAND CHANNEL ALTERNATESEACH ENGINE START .

InterfacesThe EEC receives air data parameters from the Air Data Inertional ReferenceSystem ( ADIRS ), and operational commands from the Engine Interface Unit ( EIU ) .It also provides the data outputs nescessary for the Flight Management andGuidance Computers ( FMGCs ), and the fault message to the EIU for aircraftmaintenance data system.Each EEC channel directly receives the Thrust Lever Angle (TLA ) .The EEC transmits the thrust parameters and TLA to the FMGCs for the auto-thrust function.

SensorsVarious sensors are provided for engine control and monitoring.Pressure sensors and thermocouples are provided at the aerodynamic sta-tions.The primary parameters are Engine Pressure ratio ( EPR = P4.9/P2 ), N1 andN2 speeds, Exhaust Gas Temperature ( EGT ) and metered Fuel Fuel Flow( FF ).

Fuel Metering Unit ( FMU )In the FMU, three torque motors are activated by the EEC .These provide thecorrect fuel flow , overspeed protection and Engine Shut Down.In case of an overspeed, an incorporated valve reduces the fuel flow .The fuel Pressure Raising Shut Off Valve is controlled by the EEC through theFMU , but it is closed directly from the corresponding ENG MASTER leverwhen set to OFF.

NOTE: THE FUNCTIONS OF THE FADEC ARE ALSO RESET WHEN THEENG MASTER LEVER IS SET TO OFF.

Compressor Airflow and Turbine Clearance ControlThe EEC controls the compressor airflow and the turbine clearance throughseparated sub systems.It also monitors the engine oil cooling through an air/oil heat exchanger servovalve.Compressor airflow control : Booster Stage Bleed Valves ( BSBV ). Variable Stator Vanes ( VSV ). 7th and 10th stage handling bleed valves.

Turbine clearance control : HP and LP Turbine Active Clearance Control ( ACC ) valves. 10th stage make−up air valve.(If Installed)

Engine oil cooling : Air Cooled Oil Cooler ( ACOC )servo valve .

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IDGTOIL

GEN

/PB


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FADEC DESCRIPTION

Thrust Reverser Hydraulic Control UnitThe EEC controls the thrust reverser operation through a Hydraulic ControlUnit (HCU )Each EEC channel will energize the solenoids of an isolation valve and a direc-tional valve included in the HCU to provide deployment and stowage of thethrust reverser translating sleeves.

Start and Ignition ControlEach channel can control the starter valve operation, the fuel Pressure RaisingShut − Off Valve opening and the ignition during the engine start sequence.

Fuel Diverter and Return ValveThe EEC manages the thermal exchange between the engine oil , IDG oil andengine fuel system by means of a Fuel Diverter and Return Valve.Part of the engine fuel can be recirculated to the aircraft tanks by means of areturn valve included in the fuel diverter valve module.The EEC controls the operation of the Fuel Diverter and Return Valveaccording to the engine fuel temperature ( T FUEL ) and the IDG oil tempera-ture and the engine oil temperature ( T OIL ).

Engine Parameter Transmission for Cockpit Display The FADEC provides the necessary engine parameters for cockpit displaythrough the ARINC 429 buses output.

Engine Condition Parameter TransmissionEngine Condition monitoring is provided by the ability of the FADEC totransmit the engine parameters through the ARINC 429 bus output.The basic engine parameters available are:

− WF, N1, N2, P5, PB, Pamb T4.9 (EGT), P2, T2, P3 and T3.− VSV, BSBV, 7th and 10th stage bleed commanded positions HPT/LPT ACC,HPT cooling, WF valve or actuator position− status and maintenance words, engine serial number and position.

In order to perform a better analysis of engine condition, some additional parameters are optionally available. These are P12.5, P2.5 and T2.

FADEC SYSTEM MAINTENANCE

Fault DetectionThe FADEC maintenance is facilitated by internal extensive Built in Test Equipment (BITE) providing efficient fault detection.The results of this fault detection are contained in status and maintenance words according to ARINC 429 specification and are available on the outputdata bus.

Non Volatile MemoryIn flight fault data is stored in FADEC non volatile memory and, when requested, is available on an aircraft centralized maintenance displayunit.

Communication with CFDSGround test of electrical and electronic parts is possible from cockpit,with engines not running, through the CFDS.The FADEC provides engine control system self testing to detect problemsat LRU level.FADEC is such that no engine ground run for trim purposes is necessary aftercomponent replacement.

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IDGTOIL

GEN

/PB


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FAILURES AND REDUNDANCY

Improved reliability is achieved by utilising dual sensors dual feedback.

Dual sensors are used to supply all EEC inputs exept pressures, (singlepressure transducers within the EEC provide signals to each channel−A and B ) .

The EEC uses indentical software in each of the two channels. Each chan-nel has its own power supply , processor, programme memory and input/output functions. The mode of operation and the selection of the channel incontrol is decided by the availability of input signal and output controls.

Each channel normally uses its own input signals but each channel can alsouse input signals from the other channel if required i. e. if it recognises faultyor suspect , inputs.

An output fault in one channel will cause switchover to control from theother channel.

In the event of faults in both channels a pre−determined hierarchy decideswhitch channel is more capable of control and utilises that channel.

In the event of loss of both channels, or loss of electrical power, the sys-tems are designed to go to their failsafe positions.

Single Input Signal FailureThere is no channel changeover for input signal failure, as long as the CrossChannel Data Link is operativ.

NOTE: FAULTS ARE NOT LATCHED.AUTOMATIC RECOVERY IS POSSIBLE.

Dual Input Signal FailureIf dual input signal failure occurs , the system runs on synthetized values of thehealthiest channel.The selected channel is one having the least significant failure.

Single Output Signal FailureIf an output failure occurs, there is an automatic switchover to the standby ac-tive channel.

T/S ACTION:One Channel - most likely LRU failure.

Complete output Signal FailureIn case of complete output failure there will be no current flow through torquemotors or solenoids. The associated component will be the ” FAIL−SAFE ”position.

NOTE: IF THE EEC POWER SUPPLY IS LOST, THE COMPONENTS WILLGO INTO”FAILE−SAFE” POSITION.

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TM

Page: 241Figure 122 FADEC Processing and Fault Logic

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FAILURES AND REDUNDANCYImproved reliability is achieved by utilising dual sensors dual feedback.

Dual sensors are used to supply all EEC inputs exept pressures, (singlepressure transducers within the EEC provide signals to each channel−A and B ) .

The EEC uses indentical software in each of the two channels. Each chan-nel has its own power supply , processor, programme memory and input/output functions. The mode of operation and the selection of the channel incontrol is decided by the availability of input signal and output controls.

Each channel normally uses its own input signals but each channel can alsouse input signals from the other channel if required i. e. if it recognises faultyor suspect , inputs.

An output fault in one channel will cause switchover to control from theother channel.

In the event of faults in both channels a pre−determined hierarchy decideswhitch channel is more capable of control and utilises that channel.

In the event of loss of both channels, or loss of electrical power, the sys-tems are designed to go to their failsafe positions.

ENGINE LIMITS PROTECTIONGeneralThe FADEC prevents inadvertent overboosting of the expected rating (EPRlimit and EPR target) during power setting.It also prevents exceedance of rotor speeds (N1 and N2) and burner pressurelimits. In addition, the FADEC unit monitors EGT and sends an appropriate in-dication to the cockpit display in case of exceedance of the limit.The FADEC unit also provides surge recovery.

OverspeedOverspeed protection logic consists of overspeed limiting loops, for both thelow and high speed rotors, which act directly upon the fuel flow command. Sup-plementary electronic circuitry for overspeed protection is also incorporated inthe EEC.Trip signals for hardware and software are combined to activate a torque motorwhich drives a separate overspeed valve in the fuel metering unit to reduce fuelflow to a minimum value. The engine can be shut down to reset the overspeedsystem to allow a restart if desired.

Engine surgeEngine surge is detected by a rapid decrease in burner pressure or the value ofrate of change of burner pressure, which indicates that surge varies with en-gine power level .Once detected, the EEC will reset the stator vanes by several degrees in theclosed direction, open the booster 7th and 10th stage bleeds,and lower themaximum Wf/Pb schedule.Recovery of burner pressure to its steady state level or the elapse of a timerwill release the resets on the schedules and allow the bleeds to close.

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TM

Page: 243Figure 123 FADEC Processing and Fault Logic

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POWER MANAGEMENT

Autothrust ModeThe autothrust mode is only available between idle and maximum ( MCT )when the aircraft is in flight.After take−off the lever is pulled back to the maximum climb position. The auto-thrust function will be active and will provide an EPR target for: Max climb thrust Optimum thrust An aircraft speed ( Mach number ) A minimum thrust.

Memo ModeIn the memo the thrust value is frozen to the last EPR actual value, and willremain frozen until the thrust lever is moved manually or autothrust is resetwith the autothrust pushbutton switch.When the autothrust function is disengaged while the thrust lever is in MCT/FLX or CL (Maximum Continuous / Flexible Take−Off or Climb ) detent, thethrust is locked until the thrust lever is moved manually.Memo mode or Thrust locked is entered automatically from autothrust modewhen: The EPR target is invalid, Or one of the two instinctive disconnect pushbutton switches on the thrust

levers is activated, Or autothrust signalis lost from EIU.

Manual ModeThis mode is entered any time the conditions for autothrust or memo modesare not present. In this mode, thrust lever sets an EPR value proportional tothe thrust lever position up to maximum take−off thrust.

Flexible take−off ratingFLEXIBLE TAKE−OFF rating is set by the assumed temperature method withthe possibility to insert an assumed temperature value higher than the maxi-mum one certified for engine operation . ( 30 deg C.)

AUTOTHRUST ACTIVATION / DEACTIVATION

The autothrust function (ATHR) can be engaged or active.The engagement logic is done in the Flight Management Computer (FMGC)and the activation logic is implemented into the EEC.The activation logic in the EEC unit is based upon two digital discretes:

–ATHR engaged,–ATHR active

from the FMGC, plus an analog discrete from the instinctive disconnect push-button on the throttle.The ATHR function is engaged automatically in the FMGC by auto pilotmode demand and manually by action on the ATHR pushbutton located onthe Flight Control Unit (FCU).The ATHR de−activation and ATHR disengagement are achieved by actionon the disconnect pushbutton located on the throttle levers or by depressingthe ATHR pushbutton provided that the ATHR was engaged, or by selection ofthe reverse thrust.If the Alpha Floor condition is not present, setting at least one throttle lever for-ward of the MCT gate leads to ATHR deactivation but maintains ATHR en-gaged .If the Alpha Floor condition is present, the ATHR function can be activated re-gardless of throttle position.The thrust is controlled by the throttle lever position and ATHR will be activatedagain as soon as both throttles are set at or below MCT gate.When ATHR is deactivated (pilot’s action or failure), the thrust is frozen to theactual value at the time of the deactivation. The thrust will be tied to the throttlelever position as soon as the throttles have been set out of the MCT or MCLpositions.

NOTE: AUTOTHRUST IS ONLY ACTIVE IN EPR MODE. IN RATED & UN-RATED N1 MODE AUTOTHRUST IS LOST.

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EECFMGC

FUEL FLOW

COMMANDEPR

AUTOTHRUST DISC.PB

Page: 245Figure 124 Thrust Control Architecture

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THIS PAGE INTENTIONALLY LEFT BLANK

Page: 246

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Page: 247FRA US/T Bu August 2001 Page: 247Figure 125 Auto Thrust Defenition

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Engine Fuel and ControlsFADEC Power Management

A319 / A320 / A321IAE V2530−A5

73−20


Alpha Floor ConditionIf the Alpha Floor condition is not present, setting at least one throttle lever for-ward of the MCT gate leads to ATHR deactivation but maintains ATHR en-gaged ; the thrust is controlled by the throttle lever position and ATHR will beactivated again as soon as both throttles are set at or below MCT gate.If the Alpha Floor condition is present, the ATHR function can be activated re-gardless of throttle position.When ATHR is deactivated (pilot’s action or failure), the thrust is frozen to theactual value at the time of the deactivation. The thrust will be tied to the throttlelever position as soon as the throttles have been set out of the MCT or MCLpositions.

Manual ModeThe thrust is controlled manually (i.e., function of TLA position) if the throttlesare not in the ATHR area.This mode is also entered any time the conditions for autothrust or memomodes are not present. In this mode, thrust lever sets an N1 value proportionalto the thrust lever position up to maximum take−off thrust.TLA versus rated thrust is consistent regardless of ambient conditions.TAKE−OFF/GO−AROUND ratings are always achieved at full forward throttlelever position (except in Alpha−floor mode).Other ratings (MAX CONTINUOUS, MAX CLIMB. IDLE, MAX REVERSE) areachieved at constant throttle lever positions.FLEXIBLE TAKE−OFF for a givenderating is achieved at constant retarded throttle lever position.

Flexible take−off ratingFLEXIBLE TAKE−OFF rating is set by the assumed temperature method withthe possibility to insert an assumed temperature value higher than the maxi-mum one certified for engine operation to provide for the maximum derate al-lowed by the certifying Authorities.

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Engine Fuel and ControlsFADEC Power Management

A319 / A320 / A321IAE V2530−A5

73−20


DETENTDETENT DETENT

EPR

EPR RATING

EPR TARGET

EPR REQUIRED

EPR TARGET

Page: 249Figure 126 Thrust Lever Positions

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EPR SETTING REQUIREMENTS

EPRThe EEC uses closed loop control based on EPR or, if EPR is unoptainable, onN1.Under EPR control, the EPR target is compared to the actual EPR to deter-mine the EPR error.The EPR error is converted to a rate controlled Fuel Flow command ( FF )which is summed with the measured fuel flow ( FF actual ) to produce the FFerror.The FF error is converted to a current ( I ) which is sent to the dual torque mo-tor.The torque motor repositions the Fuel Metering Valve ( FMV ) to change thefuel flow.

The inputs required for EPR control are: Ambient temperature ( T amb ) Engine air inlet temperature ( T2 ) Altitude ( ALT ) Mach number (Mn ) Throttle Resolver Angle ( TRA ). Service Bleeds

It is possible to re-select the primary control mode ( EPR) through the N1 modeP/B switch following an automatic reversion to rated or unrated N1 mode.If the fault is still present, the EEC will remain in its current thrust settingmode.If the fault is no longer present, the EEC will switch to the primary controlmode (EPR). If the fault later reoccurs,reversion back to N1 mode (rated orunrated) will result.

RATED N1 SETTING REQUIREMENTS

Rated N1The loss of either the P2 or the P 4.9 signal will cause an automatic reversionto the rated N1 closed loop control.This is a alternate control mode which utilizes to control the thrust automatical-ly.It is a despatchable mode but autothrust is not available when operating inthis mode.The rated N1 mode can also be manually selected by actuating therelated N1 MODE P/B switch (one per engine) that is located on the overheadpanel.

The inputs required for Rated N1 control are: - T2 and - the Throttle Resolver Angle ( TRA ).

The processing of the N1 error signal is the same as for EPR error signal.

UNRATED N1 SETTING REQUIREMENTS

Unrated N1The loss of the T2 signal will cause automatic reversion to unrated N1 closedloop control.Max N1,N1 thrust lever,N1 mode and N1 rating limit indications on the upperECAM are lost.

The input required for unrated N1 control is:- the Throttle Resolver Angle ( TRA).

The unrated N1 thrust setting requires the thrust to be set manually to an N1speed.An overboost can occur in the unrated N1 thrust setting at the full for-ward thrust lever position.Use of unrated N1 thrust setting overboost abovenormal rated thrust is not recommended and will result in reduced engine life.The maximum N1 must therefore be determined from charts in the Flight CrewOperating Manual ( FCOM ). It is a non-despatchable mode and autothrust is not available when operating inthis mode.The processing of the N1 error signal is the same as for the rated N1 error signal.

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Page: 251FRA US/T Bu August 2001 Page: 251Figure 127 Power Setting Requirements Schematic

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ENGINE FUEL & CONTROLFADEC POWER MANAGEMENT

A319/A320/A321IAE V2530 − A5

73−20


IDLE CONTROL

Minimum idle ( 56 % - 60% N2 ) is corrected for ambient temp >30° C, then N2 will increase.

Approach idle (approx. 70% N2 ) It varies as a function of Total Air Temperature ( TAT ) and altitude.

This idle speed is selected to ensure sufficiently short accelleration time togo around thrust and is set when the aircraft is in an approach configura-tion.(Flap Lever Position -” NOT UP”)

Reverse Idle ( approx. 70% N2 ) = Approach Idle + 1000 RPM ,FADEC sets the engine speed at reverse idle

when the throttle is set in the reverse idle detent position .

Bleed Idle = Bleed demand.Bleed Idle command will set the fuel flow requested for

ensuring correct aircraft ECS system pressurization ,wing anti ice and en-gine anti ice ,pressurization ( Pb-”ON” or valves not closed ) .

HMS Idle (Min Idle - Approach Idle) For conditions where the compensated fuel temperature is greater than 140

deg. C. , the heat management control logic calculates raised idle speed.(in flight and on ground !)

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ENGINE FUEL & CONTROLFADEC POWER MANAGEMENT

A319/A320/A321IAE V2530 − A5

73−20


ApproachIdle

BleedIdle

ReverseIdle

EEC

N2 Idle

Setting

HMS

TLA (REV. IDLE)

LGCIU1 / 2

ZONECONT.

0

1

2

3

FULL

0

1

3

FULL

2

EIUWOW (GRD)

THRUSTLEVERS

LANDINGGEARS

SLAT /FLAPLEVER SFCC

1 / 2EIU

AIR

LEVER NOT ZERO

Min Idle

EIU

WING ANTI ICE

ENG ANTI ICE

EIU FAULT

PACKs

PACKCONT.1 / 2

ECS DEMAND

ENGINE FUEL TEMPERATURE

Page: 253Figure 128 Idle Control Requirements

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ENGINE FUEL AND CONTROLFADEC POWER MANAGEMENT

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Page: 254FRA US/T Kh August 2001

N1 SPEED TABLE

RPM % N15650 100 %

5465 96,7 %

5085 90 %

4918.5 87 %

4520 80 %

4372 77,4 %

3955 70 %

3825,5 67,7 %

3390 60 %

3279 58,0 %

2825 50 %

2732,5 48,4 %

2260 40 %

2186 38,7 %

1695 30 %

1639.5 29 %

1130 20 %

1093 19,3 %

Page: 254

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ENGINE FUEL AND CONTROLFADEC POWER MANAGEMENT

A319/A320/A321IAE V2530 − A5

73−20

Page: 255FRA US/T Kh August 2001

760050,8%

800053,5%

840056,1%

880058.8%

920061,5%

960064,2%

N2

RO

TOR

SP

EE

D (

RP

M /

% )

−80 −60 −40 −20 0 +10 +20 +30 +40 +50

AMBIENT TEMPERATURE ( DEG. C. )

V2530-A5 SLS / STD GROUND IDLE ( NO OFFTAKES )

+15

57,5%

Page: 255Figure 129 Ground Idle Speed Diagram N2

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FADEC FAULT STRATEGY

GeneralThe Electronic Engine control ( EEC ) system is dual, the two channels areequal.Failures are classified as class 1, 2 , 3 .According to the failure class, the system can use data from the other channel,or switch to the other channel. Faults are memorized in the system BITE asthey occur.

Input Fault StrategyAll sensors and feedback signals are dual.Each parameter sensor as well as feedback sensors used by each channelcome from two different sourses : Local or cross− channel through the Cross channel Data Link

NOTE: SOME SENSORS CAN DIRECTLY BE SYNTHETIZED BY THECORRESPONDING CHANNEL

Single Input Signal FailureThere is no channel changeover for input signal failure, as long as the CrossChannel Data Link is operativ.

NOTE: FAULTS ARE NOT LATCHED.AUTOMATIC RECOVERY IS POSSIBLE.

Dual Input Signal FailureIf dual input signal failure occurs , the system runs on synthetized values of thehealthiest channel.The selected channel is one having the least significant failure.

Single Output Signal FailureIf an output failure occurs, there is an automatic switchover to the standby ac-tive channel.

T/S ACTION:One Channel - most likely LRU failure.

Complete output Signal FailureIn case of complete output failure there will be no current flow through torquemotors or solenoids. The associated component will be the ” FAIL−SAFE ”position.

NOTE: IF THE EEC POWER SUPPLY IS LOST, THE COMPONENTS WILLGO INTO”FAILE−SAFE” POSITION.

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.

Page: 257Figure 130 FADEC Single Input Signal Failure

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COMPONENT FAIL SAFE STATES

COMPONENTS: FAIL SAFE STATE:

METERING VALVE MIN FLOW

VARIABLE STATOR VANE ACTUATOR VANES OPEN

2.5 BLEED ACTUATOR (BSBV) BLEED OPEN

7TH STAGE HANDLING BLEED VALVES BLEED OPEN

10TH STAGE HANDLING BLEED VALVE BLEED OPEN

HPT ACC VALVE VALVE CLOSED

LPT ACC VALVE VALVE PARTIALLY OPEN - 45%

ACOC AIR VALVE OPEN

10TH STAGE ”MAKEUP” AIR VALVE OPEN

FUEL DIVERTER VALVE FMU RETURN FLOW THROUGH FCOC (MODE 4 OR 5 )SOLENOID DE-ENERGIZED

RETURN TO TANK VALVE CLOSED ( MODE 3 OR 4 )

IGNITION ON

STARTER AIR VALVE CLOSED

P2/T2 PROBE HEAT OFF

THRUST REVERSER CONTROL UNIT * REVERSER STOWED

NOTE: IF THERE IS A FAILURE OF THE THRUST REVERSERHYDRAULIC CONTROL UNIT DIRECTIONAL VALVEWHILE THE REVERSER IS DEPLOYED,THE REVERSER WILL REMAIN DEPLOYED.

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LOSS OF INPUTS FROM AIRCRAFT

EIU SIGNALS: NO ENGINE STARTING. NO AUTOTHRUST ON BOTH ENGINES.NO REVERSE THRUSTMODULATED IDLE NOT AVAILABLE.CONTINUOUS IGNITION

ADC SIGNALS: EEC USES ENGINE SENSORS.

BOTH TLA: IN REVERSE: IF REVERSER INADVERTENTLY DEPLOYS ANDBOTH REVERSER FEEDBACKS ARE INVALID,POWER IS SETTO IDLE.ON GROUND: SET IDLEIN FLIGHT: AT TAKE OFF FREEZE LAST VALID TLA,THEN SE-LECT MCT AT SLAT RETRACTIONAUTOTHRUST CAPABILITY.

ONE TLA: THE EEC USES THE REDUNDANT SENSOR.

BOTH 115V AC: NO IGNITIONNO P2/T2 PROBE HEATING

BOTH 28V DC: NO STARTRUN ON ALTERNATOR ABOVE 10% N2

DISAGREEMENT BETWEEN TRA: ON GROUND: SET FORWARD IDLEIN FLIGHT: SELECT LARGER VALUE BUT LIMIT THIS

TO MCTON REVERSE: SELECT REVERSE IDLE.

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ENGINE FUEL AND CONTROLFADEC TEST

A319/A320/A321IAE V2530 −A5

73−20


ATA 73-20 FADEC TEST

GENERAL:To get access to the FADEC SYSTEM REPORT / TEST menu the FADECGRD PWR must be switched ”ON”. Then press the line key adjacent toCFDS - SYSTEM REPORT / TEST - NEXT PAGE - ENG 1A (1B),(2A),(2B).

FADEC PREVIOUS LEGS REPORT

This CFDS menu function gives access to the faults which have been de-tected and stored during the previous 64 flight legs.The Cells indicate if the failure was detected in the ground memory or the flightmemory.

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<LAST LEG REPORT

<PREVIOUS LEG REPORT

<TROUBLESHOOTING REPORT

<SYSTEM TEST CLASS 3>

<GROUND SCANNING

<RETURN

NEXTPAGE

FADEC A FAULT

Page: 261Figure 131 Previous Legs Report

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FADEC TROUBLESHOOTING REPORT

The trouble shooting menu has 4 submenus: FLIGHT DATA GROUND DATA AIRCRAFT DATA EEC CONFIGURATION

FLIGHT DATAThis menu gives additional failure data (temperatures,pressures,RPM,etc.)when a fault occured during the flight. This data is saved in a CELL.Each CELLprovides 2 menu pages of troubleshooting informations.The cell allows a identi-fication which CFDS FAULT message belongs to which troubleshooting data(eg.Ground Scanning menu.)In the example a OSPXCF (OVERSPEED CROSS CHECK FAILURE )is indicated.

GROUND DATAThis menu gives additional failure data (temperatures,pressures,RPM,etc.)when a fault occured on ground. This data is saved in a CELL.The cell allows a identification which CFDS FAULT message belongs to whichtroubleshooting data (eg.Ground Scanning menu.)

FADEC FAILURE TYPES DEFINITION

WRAP - AROUND FAILURE (WAF)A detected failure in the circuitry of a system.The EEC checks for continuity.If failed in one channel: EEC switches to the other channel

(the ability to switch is based on relative helth of the other channel)If failed in both channels: specific output is depowered

(exception - solenoids are depowered in groups)T/S ACTION:Most likley a loose connector or chaffed harness next LRU and finally EEC.

TRACK-CHECK FAILURES (TKF)Failure of the system to follow the commands of the EEC.The EEC compares feedback position against commanded position.If failed in one channel: EEC switches to the other channel

(the ability to switch is based on relative helth of the other channel)If failed in both channels: Healthiest channel continues to command actuator.

T/S ACTION:one channel - most likely LRU failure.both channels - most likely mechanical failure ,check LRU/moving mechanism.

CROSS CHECK FAILURES (XCF)A detected difference in the feedbacks from the LRU LVDT‘s or microswitches.The EEC compares channel A against Channel B.Failure of TRA: EEC has specific fault accomodation based on previous value.Failure of Reverser: EEC will select most stowed and will not allow a deploy.Failure of Temperature sensors: EEC will use fail safe value.T/S ACTION:Most likely a LRU problem ,next check harness then EEC

INPUT LATCHED FAILED (ILF)(Single Input Signal Failure )There is no channel changeover for input signal failure, as long as the CrossChannel Data Link is operativ.

NOTE: FAULTS ARE NOT LATCHED. THUS AUTOMATIC RECOVERY ISPOSSIBLE.

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TROUBLE SHOOTING

<FLIGHT DATA

<GROUND DATA

<AIRCRAFT DATA

<EEC CONFIGURATION

<RETURN

<LAST LEG REPORT




<GROUND SCANNING

<RETURN

A second page is available to givemore trouble shooting data

Page: 263Figure 132 Trouble Shooting Report

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FADEC 1BFLIGHT DATA

CELL: 31 FAULT: WOFWAF

PG:01

RPM: N1 = 5326 N2 = 14392DEG C: T5 = 554.0 T2 = 26.0TCJC = 42.0 FLTPH = 3

PSIA: PB = 458.5 P2 = 14.62

MN = .117 HOURS = 571.0

FADEC FAULT CELL

N1 RPM

T5 Temperature( T4.9 EGT )

Cold Junction Temperature( Actual Temp. in EEC )

Air Pressure on Eng. Station 3( PB = Burner Pressure )

Mach Number

Fault Code

N2 RPM

T2 Temperature ( Eng. Inlet )

Flight Phase

Total Air Pressure ( Eng. Station 2 )

EEC Operating Hours

NOTE: THE ABBREVIATIONS USED IN THE GROUND DATA ARE THE SAME.

Page one of the Cell 31

Page: 264Figure 133 Flight Data / Ground Data

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FADEC 1BFLIGHT DATA

CELL: 31 FAULT: WOFWAF

PG:02

ALT: = 336.0 FT EPRI = 1.562SVA : = 1.906 INCH INCOM = 1FF = 11162 PPH BACKUP = 0B 25 = 1.218 INCH LEG = 398.0

WOW = 1

FADEC Fault Cell

Standart Altitude

Stator Vane Actuator( Feedback )

Fuel Flow

2.5 Bleed Actuator Feedback

Weight on Wheels1 = Yes ( Ground )0 = NO ( Flight )

Fault Code

EPR ( indicated )

Channel in Control1 = Yes , 0 = NoN1 Mode1 = Yes0 = No ( EPR Mode )

Flight Legs

NOTE: THE ABBREVIATIONS USED IN THE GROUND DATA ARE THE SAME.

Page two of the Cell 31

Page: 265Figure 134 Flight Data / Ground Data

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FADEC SYSTEM TEST

The FADEC SELF TEST should really be known as the FADEC SYSTEMTEST.The test and results can be split into three categories described as follows.

Output Driver Test This is a systen maintenance test that performs a wraparound (continuity) testof all the EEC output driver lines and associated component wiring.There are three possible results as follows:1. Output Driver Test Failed − Indicates that a continuity fault was found.2. Output Driver Test Passed − Indicates that no wraparound fault was

found.3. Output Driver Test No Run − Indicates that the test was not run because

the tested channel was not capable of powering the outputs.

Input / lnternal TestThis is the FADEC (EEC) internal check to verify that the local channel interface, input and output circuits are functional prior to entering MENUMODE. There are three possible results as follows:4. Input / Internal Test Failed − Indicates that the activity monitor circuit test

failed or the local channel was unable to provide power to any Output orthere were interface or input fault.

5. Input / lnternal Test Passed − Indicates that the activity monitor circuitpassed and that no interface or input faults were set prior to entry intomenu mode.

6. Input / Internal Test No Run − Indicates that the local cannel was not capable of powering its outputs or that the EEC has not spent the minimumof 30 seconds in normal mode.

Pressure Sensor TestThis is an internal measurement of the pressure sensors (P2, P5, Pb, PMX) inthe EEC via the local channel to make sure they are within a specified tolerance of each other.The three possible results are as follows:7. Pressure Sensor(s) Failed − Indicates that an interface or range failure

(from normal mode) is set for any pressure sensor (hard failures).8. Pressure Sensor(s) Agree − Indicates that the static pressure sensor test

ran and that all the pressure sensors are within tolerances.9. Pressure Sensor(s) Disagree − Indicates that the static pressure sensor

test ran and any two pressure sensors were not within the specified tolerances.

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<LAST LEG REPORT




<GROUND SCANNING

<RETURN

SYSTEM TEST

<FADEC SELF TEST

<REVERSER TEST

<IGNITOR TEST

<STARTER VALVE TEST

<RETURN

NOTE: IF EVERY TEST FAILED,RETURNTO FADEC / MENU PUSH THELINE KEY ADJACENT TOGROUND SCANNING ANDCHECK THE FAILURE MESSAGE.

Page: 267Figure 135 FADEC Self Test

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A319/A320/A321IAE V2530 −A5

73−20


FADEC GROUND SCANNINGThis menu shows the faults which are present on ground.More information canbe obtained using the troubleshooting menu.This menu must also be used to indicate which faults were detected in theother FADEC TEST menus (eg. Starter Valve Test,Reverser Test,etc.)

FADEC CLASS 3 FAULT REPORTThis menu shows all class 3 faults of the FADEC system which have to re-paired after 200 hours or during an A-maintenance check.

Page: 268

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73−20


<LAST LEG REPORT




<GROUND SCANNING

<RETURN

CLASS 3 FAULT

Page: 269Figure 136 Ground Scanning

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A319/A320/A321IAE V2530 −A5

73−20


FADEC CLASS 3 FAULT REPORTThis menu shows all class 3 faults of the FADEC system which have to re-paired after 200 hours or during an A-maintenance check.

Page: 270

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<LAST LEG REPORT




<GROUND SCANNING

<RETURN

CLASS 3 FAULT

Page: 271Figure 137 FADEC Class 3 Fault Report

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ENGINE FUEL AND CONTROLENGINE INTERFACE UNIT

A319/A320/A321IAE V2530 A5

73−20


ATA 73-25 ENGINE INTERFACE UNIT

EIU PRESENTATION

Two EIUs are fitted on each aircraft, one for engine 1, one for engine 2Each EIU, located in the electronics bay 80VU, is an interface concentrator between the airframe and the corresponding FADEC located on the engine,thus reducing the number of wires. EIUs are active at least from engine startingto engine shutdown, they are essential to start the engine.

The main functions of the EIU are:− to concentrate data from cockpit panels and different electronic boxes to the associated FADEC on each engine,− to insure the segregation of the two engines,− to select the airframe electrical supplies for the FADEC,− to give to the airframe the necessary logic and information from engine to other systems (APU, ECS, Bleed Air, Maintenance).

EIU INPUT DESCRIPTION

EIU input from the EECThe EIU acquires two ARINC 429 output data buses from the associated EEC(one from each channel) and it reads data from the channel in control. Whensome data are not available on the channel in control, data from the otherchannel are used.In the case where EIU is not able to identify the channel in control, it will as-sume Channel A as in control.The EIU looks at particular engine data on the EEC digital data flow to interfacethem with other aircraft computers and with engine cockpit panels.

EIU output to the EECThrough its output ARINC 429 data bus, the EIU transmits data coming from allthe A/C computers which have to communicate with the EEC, except fromADCs and throttle which communicate directly with the EEC.There is no data flow during EIU internal test or initialization. EIU Location

824

80VU

EIU

Page: 272

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A319/A320/A321IAE V2530 A5

73−20


To CIDS (23−73 )

To DFDRS INTCON Monitoring (31−33)

To CVR power Supply (23−71 )

To Avionics Equipment Ventilation (21−26 )

To WHC (30−42 )

To PHC ( 30−31 )

To FCDC (27−95)

To Blue Main Hydraulik PWR( 29−12)

To Green Main HYD PWR RSVR Indicating (29−11)

To Yellow Main HYD PWR RSVR Indicating (29−13 )

To Blue Main HYD PWR RSVR Warning / Indicating

Page: 273Figure 138 EIU Schematic

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73−20


EIU INTERFACES

SIGNALS PURPOSE

WING ANTI-ICE SWITCH ENGINE BLEED COMPUTATION LOCIG

ENGINE FIRE P/B SIGNAL FADEC ENGINE SHUTDOWN LOGIC

LOW OIL PRESSURE SWITCH (AND GROUND) -COCKPIT WARNING SIGNALS-HYDRAULIC MONITORING-WINDOW AND PROBE HEATING SYSTEM-AVIONIC VENTILATION SYSTEM-RAIN REPELLENT SYSTEM-CIDS,CVR,DFDR

FADEC GROUND POWER P/B FADEC POWER SUPPLY LOGIC

LGCIU 1 AND 2 (GROUND SIGNAL) THRUST REVERSER AND IDLE LOGIG

SFCC 1 AND 2 ENGINE FLIGHT IDLE COMPUTATION LOGIC

SEC 1 ,2 AND 3 THRUST REVERSER INHIBITION CONTROL

FLSCU 1 AND 2 HEAT MANAGEMENT SYSTEM FUEL RETURN VALVECONTROL

ENGINE SELECTED ENGINE 1 OR 2 INDENTIFICATION

OIL PRESSURE,OIL QUANTITY AND OIL TEMPERATURE INDICATION ECAM

NACELLE TEMPERATURE INDICATING (ECAM)

START VALVE POSITION (FROM EEC) ECS FOR AUTOMATIC PACK VALVE CLOSURE, DURINGENGINE START

N2 GREATER THAN MINIMUM IDLE (FROM EEC) FUNCTIONAL TEST INHIBITION OF THE RADIO ALTIME-TER TRANSCEIVER -BLUE HYDRAULIC SYSTEM PUMP CONTROL

ENGINE START FAULT SIGNAL ILLUMINATION OF FAULT LIGHT ON THE ENGINE STARTPANEL

APU BOOST DEMAND SIGNAL (EIU) MAIN ENGINE START MODE TO THE APU ELECTRONICCONTROL BOX

Page: 274

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73−20


EIU INTERFACES CONT.

SIGNALS PURPOSE

TLA IN TAKE-OFF POSITION (MIN. T/O N2, FROM EEC) PACK CONTROLLER FOR INLET FLAP CLOSURE-AVIONIC EQUIPMENT VENTILATION CONTROLLER( CLOSED CIRCUIT CONFIGURATION )-CABIN PRESSURIZATION COMPUTER PRE-PRESSUR-IZATION MODE

THRUST REVERSER (FROM SEC 1,2 AND 3 ) THRUST REVERSER INHIBITION RELAY

Page: 275

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A319/A320/A321IAE V2530 A5

73−20


CFDS SYSTEM REPORT/TEST EIUThis Page shows the menu of the Engine Interface Unit ( EIU )The EIU is a Type 1 System. The EIU is availlable in CFDS back up Mode.

Page: 276

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A319/A320/A321IAE V2530 A5

73−20

Page: 277FRA US/T Bu August 2001 Page: 277Figure 139 EIU Menu

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A319/A320/A321IAE V2530 A5

73−20


LAST LEG REPORT

Last leg ReportHere are Displayed the Internal EIU Faillures that Occured during Last Flights.

LRU INDENTIFICATIONShows the EIU part number.

Page: 278

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A319/A320/A321IAE V2530 A5

73−20

Page: 279FRA US/T Bu August 2001 Page: 279Figure 140 Last Leg Rep./ LRU Indentification

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GROUND SCANNINGThis Page gives the EIU Faillures still presend on Ground. RTOK means Re - Test Ok, you can ignore this Fault

Page: 280

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73−20

Page: 281FRA US/T Bu August 2001 Page: 281Figure 141 Ground Scanning

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A319/A320/A321IAE V2530 A5

73−20


EIU CFDS DISCRETE OUTPUTS SIMULATION

The Purpose of this Menu is to Simulate some Engine Interface Unit ( EIU )Discrete Outputs by Setting their Status to 0 or1 .

WARNING:

The DISCRETE OUTPUT SIMULATION can operate systems and compo-nents without special indication on the MCDU. Make allways sure that theworking areas are clear ! For the simulation refer to AMM 73-25-34 , (TASK 73-25-34-860-041).The Discrete Outputs are Listed on two Pages, one for the Positive Type andone for the Negative Type.

SIMULATION : ” APU BOOST ”To simulate an APU BOOST command through the MCDU.Push the line key adjacent to”APU BOOST” discrete output status:”APU BOOST”becomes ”1” and the EIU sends the APU BOOST command tothe 59KD ECB.APU BOOST 1 simulates a not closed starter air valve.The APU is boosted(if running)APU BOOST 2 simulates a energized starter air valve solenoid.

SIMULATION : ” FAULT ”To simulate a disagree between the position and the command of the HP fuelvalve through the MCDU the line key adjacent toFAULT discrete output statusismust be pushed.The FAULT becomes ”1” and the FAULT legend of the5KS1(2) annunciator light comes on.

SIMULATION : ”LOP GND 1 ”To simulate ”OIL LOW PRESS & GND” for the following systems through theMCDU :PHC1, PHC3, WHC1, AEVC, DFDR and CVR.CAUTION : REMOVE THE PROTECTIVE COVERS FROM THE PROBESBEFORE YOU DO THE TEST.If the line key adjacent to LOP is pused, LOP GND1 discrete output status be-comes GND1 ”0”

The PHC1(3) commands a low probes heating levelThe WHC1 commands a low captain windshield heating levelThe CVR and DFDR are switched onNOTE : When ”LOP GND1” is simulated to ”0” the horn will be inhibited incaseof low avionic bay extract airflow.

SIMULATION : ”LOP GND 2 ”To simulate ”OIL LOW PRESS & GND” for the following systems through theMCDU :Blue / yellow main hydraulic pressure power warning indicating WHC2,PHC2, green main hydraulic PWR RVSR indicating, FCDC1, FCDC2.When the line key adjacent to LOP ”LOP GND2 ” discrete output status be-comes GND2 ”0”.”B (Y) ELEC PUMP LO PR” warning message is no longer inhibited The PHC2 commands a low probes heating levelThe WHC2 commands a low windshield (F/O) heating levelThe 3DB1 and 3DB2 rain repellant valve opening is authorized

NOTE: THE ”LOP GND2” DISCRETE IS USED TO INHIBIT THE FLIGHTCONTROL SYSTEM TEST THROUGH THE CFDS. ACCESS TOTHIS MENU IS PROHIBITED BY THE CFDS ARCHITECTURE ASLONG AS YOU WORK ON THE EIU DISCRETE OUTPUTS MENU.

SIMULATION : ” T/R INHIB ”To simulate the authorization of the T/R directional control valve solenoid clo-sure (through the 14KS1(2) relay) through the MCDU.When the line key adjacent to T/R is pushed, ” T/R INHIB ” discrete outputstatus INHIB becomes ”1” and the 14KS1(2) inhibition relay is energized, au-thorizing the directional control valve solenoid energization

Page: 282

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73−20


APU BST2

APU BST1

Page: 283Figure 142 Discrete Outputs Simulation

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A319/A320/A321IAE V2530 A5

73−20


EIU CFDS DISCRETE OUTPUTS SIMULATION

SIMULATION : ” HP FUEL PN ”To simulate a HP FUEL VALVE 1(2) in open position through the MCDU.Push the line key adjacent to HP”HP FUEL PN” discrete output statusFUEL PN becomes ”1” and the zone controller 8HK will receive the HP FUELVALVE 1(2) open condition.

NOTE: THE ZONE CONTROLLER USES THE HP FUEL VALVE POSITIONTO ELABORATE THE BLEED STATUS ON LABEL 061 AND SENDSIT TO THE EEC THROUGH THE EIU (LABEL 030). THE BLEEDSTATUS CAN ONLY BE MODIFIED BY THIS INPUT IF THE PRVOPENS (ENGINE RUNNING).

SIMULATION OF ” PACKS OFF ”To simulate the PACK FLOW control valve closure command through theMCDU push the line key adjacent to”PACKS OFF” discrete output status.PACKS OFF becomes ”1” and the PACK FLOW control valve closure solenoidis energized.

NOTE: THE PACK FLOW CONTROL VALVE 1(2) REQUIRE A MUSCLEAIR PRESSURE TO OPEN.

SIMULATION OF ” N2 > IDLE ”To simulate ”N2 > IDLE” for the following systems :XCVR radio altimeter 25ABlue main hydraulic power

WARNING: MAKE SURE THAT THE TRAVEL RANGES OF THE FLIGHTCONTROL SURFACES ARE CLEAR BEFORE YOU PRES-SURIZE / DEPRESSURIZE A HYDRAULIC SYSTEM.

Push the line key adjacent to N2 . N2 > IDLE DISCRETE OUTPUT becomes”1”> IDLE The electric pump of the blue hydraulic system start and the blue hydraulic sys-tem is pressurized (approximately 3000PSI)

NOTE: THE N2 > IDLE DISCRETE IS USED TO INHIBIT THE ”RAMPTEST” OF THE RADIO ALTIMETER 1(2). ACCESS TO RADIO AL-TIMETER RAMP TEST MENU IS PROHIBITED BY THE CFDS AR-CHITECTURE AS LONG AS YOU WORK ON THE EIU DISCRETEOUTPUTSMENU.

SIMULATION OF ” TLA > MCT ”To simulate ”TLA > MCT” for the following systems :AEVC, PACK CONTROLLERSCABIN PRESSURE CONTROLLERS.Push the line key adjacent to TLA ”TLA > MCT” discrete output status > MCTbecomes ”1”On the ECAM PRESS page check that the inlet and extract skin air valvesclose .

Page: 284

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73−20


APU BST2

APU BST1

Page: 285Figure 143 Discrete Outputs Simulation

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73−20


EIU DISCRETE OUTPUTSMany systems get the engine ”on” or ”off” signal.This signal is switched via theOil Low Press and Ground relay.The relay is directley triggert from the EIU.

Low Oil Pressure Switching via EIU To CIDS (23−73 ) To DFDRS INTCON Monitoring (31−33) To CVR power Supply (23−71 ) To Avionics Equipment Ventilation (21−26 ) To WHC (30−42 ) To PHC ( 30−31 ) To FCDC (27−95) To Blue Main Hydraulik PWR( 29−12) To Valve Rain RPLNT. (30−45 ) To Green Main HYD PWR RSVR Indicating (29−11) To Yellow Main HYD PWR RSVR Indicating (29−13 ) To Blue Main HYD PWR RSVR Warning / Indicating (29−12)

Page: 286

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73−20

Page: 287FRA US/T Bu August 2001 Page: 287Figure 144 EIU Discrete Outputs

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EIU DISCRETE OUTPUTS

Page: 288

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73−20

Page: 289FRA US/T Bu August 2001 Page: 289Figure 145 EIU Discrete Outputs

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ENGINE AIRAIR SYSTEMS GENERAL

A319/A320/A321IAE V2530−A5

75−00

Page: 290FRA US/T BU August 2001

ATA 75 ENGINE AIR

75−00 SYSTEM PRESENTATION

General− Nacelle Compartement and Accessory Cooling− Bearing Compartment Cooling and Sealing− HP TurbineCooling− HP / LP Turbine Clearance Control System ( ACC )− Ignition System Cooling ( REF, ATA 74 )

75−30 Compressor Control− LP Compressor Airflow Control System− HP Compressor Airflow Control System

75−40 Nacelle Temperature Indicating

The external air system consits of the following subsystems: Fuel control system air bleed HP / LP turbine active clearance control High energy igniter harness cooling air Engine bleed air.

The internal air system consits of : Propulsion airflow ( secondary & primary flows ) Bearing compartments pressurizing air Cooling air

FADEC Compressor and Clearance Control

GeneralThe engine compressor and clearance control system are provided with servovalves operated by fuel pressure, but the HP compressor handling bleed valvesare operated by pneumatic pressure.The actuators have two feedback signals, one for channel A one for channelB, exept for the HP compressor handling bleed valves which do not have anyposition feedback.There is a cross−talk between the two channels, so that each channel knowsthe position sensed by the other channel.

Compressor and Clearance Control LRU‘s BSBV Master Actuator

− Servo Valve− Feedback for EEC

BSBV Slave Actuator− Servo Valve− Feedback for EEC

VSV Actuator− Sevo Valve− Feedback for EEC

7TH Stage Bleed Valves ( 3 )− 7th Stage Solenoids ( 3 )

10th Stage Bleed Valve− 10th Stage Solenoid

Page: 290

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ENGINE AIRAIR SYSTEMS GENERAL

A319/A320/A321IAE V2530−A5

75−00

Page: 291FRA US/T BU August 2001

BLEED VALVE (1x)

BLEED VALVE (3x)

ENGINEBLEED VALVE

FAN AIR

FAN AIRENGINEBLEED VALVE

HPC

HPC

HPT AIR VALVE

LPC BLEED

VSV

LP Turbine Active Clearance

HP Turbine Active Clearance

ENGINE STABILITY BLEED PART

LOCATIONS LOCATIONS

LOCATIONS

Page: 291Figure 146 Air Systems Schematic

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AIRGENERAL

A319/A320/A321IAE V2530−A5

75−00


TURBINE COOLING CONTROLThe EEC controls the actuation of an Active Clearance Control ( ACC ) valvefor the HP and Lp turbine active clearance control and a 10th stage make-upair valve for supplementary internal cooling of the turbines.

HPT/LPT Active Clearance Control ( HPT/LPT ACC )The active clearance control ( ACC ) system ensures the blade tip clearancesof the turbines for better performance.The HPT / LPT ACC valve modulates fan air flow to the HP and LP turbinecases.The EEC controls the valve position as a function of thrust level and altitude.The LVDT’s transmit the valve position to the EEC.

HP Turbine ( 10th Stage ) Cooling Air Control (if installed)The HP turbine cooling air valve ( make up air valve ) supplies supplementalair ( from HPcompressor 10th stage ) to cool the 2nd stage vanes, hubs anddiscs of the HP .The valve operates as a function of high rotor speed and altitude and incorporates a 2 − position switch to provide a feedback signal to the EEC ( channels A and B ). During cruise the valve is closed.

OPERATING SCHEDULEThe graph shown below shows control valve position, and actuator positionrelated to operation points A to E.

Engine StoppedWith the engine stopped, the position of the actuator piston is point A. At thispoint : The control valve for the HP turbine ACC is closed. The control valve for the LP turbine ACC is not less than 44 per cent

opened.

Engine operationDuring engine operation, the EEC controls the position of the actuator pistonbetween point B and point E.

Take − offDuring take − off, the position of the actuator piston is at point C.At this point : The control valve for the HP turbine ACC is closed. The control valve for the LP turbine ACC is not less than 70 per cent

opened.

NOTE: THE ACTUATOR POSITION BETWEEN POINT C AND POINT EDEPENDS ON ALTITUDE.

Fail SafeWhen there is no torque motor current or no fuel servo pressure, the ac-tuator piston moves to point A. LP valve will be partially open ( -44 deg )The actuator piston remains at this point at all defective conditions.( HP valve closed )

Page: 292

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AIRGENERAL

A319/A320/A321IAE V2530−A5

75−00


10TH STAGEMAKE UP AIRVALVE

HPT AND LPTCOOLING

MANIFOLDS

HPT 2NDVANES INTERNALCOOLING

10TH STAGEHP COMP AIR

TONo 4 BEARINGSCAVENCE VLV

MECHANICALLINKAGE

ACC VALVE

SERVOPRESS

LP PRESSRETURN

FMU

HPT

LPT

P3

HPT / LPTACCACTUATOR

EEC

MAKE UP SOLCONTR VLV

F / B

FAN AIR

Page: 293Figure 147 Turbine Cooling Control Schematic

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AIRGENERAL

A319/A320/A321IAE V2530−A5

75−00


HPT / LPT ACTIVE CLEARANCE CONT. SYS.The HP / LP Turbine Active Clearance Control ( ACC ) system uses fan airto cool the HP and LP cases for blade tip clearance control in order to im-prove engine performance and maximize the turbine cases life time. Fan air isdrawn from a common HP / LP turbine ACC air scoop in the fan duct. This airis divided into HP and LP cooling air and passes through individual short ductsto the Active Clearance Control Valves which direct air for both HP and LP tur-bine case cooling.

NOTE: THE HP TURBINE CLEARANCE CONTROL VALVE IS EQUIPPEDWITH 4 PLUGS IN THE VALVE VANE. THIS PLUGS CAN BE RE-MOVED ACCORDING TO A SERVICE BULLETIN TO ALLOW APERMANENT COOLING OF THE HP TURBINE.IN CASE OF A VALVE REMOVAL / INSTALLATION THE SAMECONFIGURATION MUST BE PROVIDED ON THE NEW VALVE.IF THE PLUGS MUST BE REMOVED, THERE IS A STORAGEBRACKET PROVIDED ON THE ACTUATOR ROD. DO NOTTHROW THE PLUGS AWAY !

HPT / LPT COOLING MANIFOLDS

HP Turbine ManifoldThe assembly consists of a left and right hand tube assemblies which are asimple push fit into the manifold.Air outlet holes on the inner face of the tubes direct the air onto the HP turbinecasings.

LP Turbine ManifoldThe assembly consists of a upper and lower tube assemblies with integralmanifolds, both ends of the cooling tubes are sealed.Air outlet holes on the inner surfaces direct the air onto the LP turbine cases.

HP COOLING

LP COOLING

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AIRGENERAL

A319/A320/A321IAE V2530−A5

75−00


LPT / HPT ACTVE CLEARANCE CONTROLVALVE ( ACC VALVE )

HP VALVE VANEVANE STEM

REMOVABLE PLUGS

LP

PLUGSSTORAGEBRACKET

HP VALVE

VALVE

Page: 295Figure 148 LPT / HPT Active Clearance Control Valve

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AIRGENERAL

A319/A320/A321IAE V2530−A5

75−00


HPT / LPT COOLING MANIFOLDS

HP Turbine ManifoldThe assembly consists of a left and right hand tube assemblies which are asimple push fit into the manifold.Air outlet holes on the inner face of the tubes direct the air onto the HP turbinecasings.

LP Turbine ManifoldThe assembly consists of a upper and lower tube assemblies with integralmanifolds, both ends of the cooling tubes are sealed.Air outlet holes on the inner surfaces direct the air onto the LP turbine cases.

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AIRGENERAL

A319/A320/A321IAE V2530−A5

75−00


HP COOLING

LP COOLING

Page: 297Figure 149 HPT / LPT Cooling Manifolds

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AIRGENERAL

A319/A320/A321IAE V2530−A5

75−00


COMPRESSOR CONTROL

General The booster stage bleed valve, the variable stator vane and HP compres-sor bleed valves systems are controlled by the EEC. The booster stage bleedvalve controls the LP compressor airflow. The variable stator vane and the 7thand 10th stage bleed valves control the HP compressor airflow.

Booster Stage Bleed Valve ( BSBV ) ControlThe BSBV position is controlled by the EEC. The EEC uses the BSBV feed-back signal from the LVDT to adjust the actual BSBV position.At low LP spool speeds the booster provides more air than the core engine canutilize. To match the booster discharge airflow to the core engine requirementsat low speed, excess air is bled off through booster stage bleed valves ( BSBV ) into the fan discharge air stream. At higher engine speeds the BSBVare closed so that all the booster discharge ( primary air flow ) enters the coreengine.

Variable Stator Vane ( VSV ) ControlThe VSV position is controlled by the EEC The EEC uses the VSV feedback signal from the LVDT‘s to adjust the actualVSV position.The VSV system maintains a satisfactory compressor performance over a widerange of operating conditions. The system varies the angle of the inlet guidevanes and stator vanes to aerodynamically match the low pressure stages ofcompression with the high pressure stages. This variation of vane positionchanges the effective angle at which the air flows across the compres-sor blades and vanes. The VSV angle determines the compression characteris-tics ( direction and velocity ) for any particular stage at compression.

HP Compressor Bleed Valves The 7th and 10th stages bleed valves maintain a more stable operation of thecompressor.

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AIRGENERAL

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75−00


10TH

7TH3

MASTER

3 X

X

SLAVE

Page: 299Figure 150 Compressor Control Schematic

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AIRGENERAL

A319/A320/A321IAE V2530−A5

75−00


ATA 75-31 LP COMP.AIR FLOW SYS.

BOOSTER BLEED SYSTEM

GeneralThe primary function of the LP compressor airflow control system is to control the airflow thus ensuring compressor stable operation during :

− Engine start.− Engine transient operation.

Description

GeneralThe airflow control system includes :

1.Two bleed−valve actuating rods2.Pisten Jack Fork End3.An LPC bleed−master actuator4.An LPC bleed−slave actuator5.Intermediate Structure

A booster bleed valve and actuating mechanismThe airflow control system automatically operates to control the air bled fromthe LP compressor.The two actuators are mechanically attached to each actuating rod and, thebleed − valve and actuating mechanism. The two actuators are connected hy-draulically and operate together by command and feedback signals from/ to theEEC.

FAIL SAFE POSITION : ” BSBV OPEN ”

In case of a malfunction ” ENG 1 ( 2 ) COMPRESSOR VANE ” is displayedon the ECAM E / WD.

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LVDT

EEC

Page: 301Figure 151 Booster Stage Bleed Valve System

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BSBV ACTUATING MECHANISM

Booster Bleed Valve and Actuating MechanismDescription The bleed valve and actuating mechanism is a sub − assembly which in-cludes :− The support ring. − The ring valve − The two upper arms, the lower arms and the eight mid arms. − The two actuating rods connect the two upper power arms to the two

actuators.The bleed valve and actuating mechanism operates to make each bleed valvesynchronized, in relation to the positions of the two actuators.

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Page: 303FRA US/T Bu August 2001 Page: 303Figure 152 BSBV and Actuating Mechanism

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ATA 75-32 HP COMP. AIR FLOW SYS.VSV SYSTEM COMPONENTS

The four stages of variable incidence stators comprise inlet guide vanes tostage 3 and stages 3, 4 and 5 stator vanes.

General The purpose of this system is to position the Inlet Guide Vanes ( IGV ) andstator vanes, using a fuel driven hydraulic actuator, in response to electricalsignals provided by the EEC.

Variable Stator Vane ( VSV ) ControlThe VSV position is controlled by the EEC as a function of N2 / square rootof theta T 2.6 ( synteziesed value ).The EEC uses the VSV feedback signal from the LVDT‘s to adjust the actualVSV position.

Description

Variable Stator Vane Actuator The stator vane actuator accurately controls vane movement with respect to atorque motor current supplied by the EEC. Operation of the stator vanes in reg-ulated by accurate control of high pressure fuel flow to one or other side of adifferential area piston. The piston has an externally adjustable low speed stopat the extended end of its travel. The high speed stop is formed by a collarwhich limits piston retraction.Provision is made to lock the piston with a rigging pin for setting pur-poses.

Linear Variable Differential Transformer ( LVDT ) A Dual Wound Linear Variable Differential Transformer ( LVDT ) is located inthe center of the actuator piston rod . The LVDT completes the electronic con-trol loop by providing a signal of actuator position to the Engine ElectronicControl.

Engine Linkage with the VSV Actuator The engine IGV and Stator Vane linkage is connected to a fork end on the pis-ton rod of the VSVA unit. The securing pin of link on to fork end.

Operation of the VSV Actuator Dual wound torque motors convert electrically isolated drive signals from eachchannel of the Electronics Engine Control ( EEC ) into hydraulic drive signalsto position the actuator piston. If power to the stator vane actuator torque motor is lost, the stator vaneactuator will go to the full open position.

Variable Stator Vane Actuation Mechanism The variable geometry operating mechanism for the compressor comprises thefollowing elements

− actuator/crankshaft drag link − crankshaft (steel) − four crankshaft/unison ring drag links − four unison rings − spindle levers ( titanium ) − variable IGVs and stage 3, 4, and 5 variable stators

FAIL SAFE POSITION : ” VANES OPEN ”

In case of a malfunction ” ENG 1 ( 2 ) COMPRESSOR VANE ” is displayedon the ECAM E / WD.

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RIG HOLES

ACTUATOR

Page: 305Figure 153 VSV System Components

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VSV RIGGING

Variable Stator Vane System ( VSVS )

Actuator Installation / RiggingBefore the actuator is removed it is important that the VSV crankshaft assembly is locked in order to prevent damage to the stator vanes.Rig pins are provided to lock the crankshaft and the actuator, as shown below.After the fuel supply and return tubes have been disconnected the crankshaftshould be rotated to align the rig pin holes in the input lever and the front bearing housing.Spanner ( Wrench ) flats are provided on the crankshaft for this purpose.Installing the rig pin locks the crankshaft assembly with the actuator and vanesin the high speed position ( actuator fully retracted ).

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L/H

R/H

Page: 307Figure 154 VSV Actuator Rig

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HANDLING BLEED VALVESHandling bleed valves are fitted to the HP compressor to improve engine start-ing, and prevent engine surge when the compressor is operating at off−designconditions.A total of four bleed valves are used, three on stage 7 and one on stage10.The handling bleed valves are two position only − fully open or fully closed,and are operated pneumatically by their respective solenoid control valve.The solenoid control valves are scheduled by the EEC.When the bleed valves are open air bleeds into the f an duct through ports inthe inner barrel of the ” C ” ducts.The servo air used to operate the bleed valves is HP compressor delivery airknown as P3 or Pb.Silencers are used on some bleed valves.All the bleed valves are spring loaded to the open position and so will alwaysbe in the correct position ( open ) for starting.

DescriptionThe bleed valve is a two position valve and is either fully open or fully closed.The bleed valve is spring loaded to the open position and so all the bleedvalves will be in the correct position - open - for the engine start. When the engine is started the bleed air from the engine will try to close the valve. Thevalve is kept in the open position by servo air ( P3 ) supplied from the sole-noid control valve ( solenoid de-energised ). The bleed valves will be closed atthe correct time during an engine acceleration by the EEC energising the sole-noid control valve vents the P3 servo air from the opening chamber of thebleed valve, and the bleed valve will move to the closed position.

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7 A

NOTE: FOR TROUBLE SHOOTING PUR-POSES A PNEUMATIC TESTSET ISAVAILABLE TO TEST THE OPERA-TION OF THE BLEED VALVES,BE-CAUSE ONLY THE SOLENOIDVALVES ARE MONITORED !

Page: 309Figure 155 HP Compressor Bleed Valves

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HANDLING BLEED VALVESHandling bleed valves are fitted to the HP compressor to improve engine start-ing, and prevent engine surge when the compressor is operating at off−designconditions.A total of four bleed valves are used, three on stage 7 and one on stage10.The handling bleed valves are two position only − fully open or fully closed,and are operated pneumatically by their respective solenoid control valve.The solenoid control valves are scheduled by the EEC.When the bleed valves are open air bleeds into the f an duct through ports inthe inner barrel of the ” C ” ducts.The servo air used to operate the bleed valves is HP compressor delivery airknown as P3 or Pb.Silencers are used on some bleed valves.All the bleed valves are spring loaded to the open position and so will alwaysbe in the correct position ( open ) for starting.

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75−00


CCDL

Page: 311Figure 156 HP Compressor Bleed Valves

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HANDLING BLEED VALVES FUNKTION

DescriptionThe bleed valve is a two position valve and is either fully open or fully closed.The bleed valve is spring loaded to the open position and so all the bleedvalves will be in the correct position - open - for the engine start. When the engine is started the bleed air from the engine will try to close the valve. Thevalve is kept in the open position by servo air ( P3 ) supplied from the sole-noid control valve ( solenoid de-energised ). The bleed valves will be closed atthe correct time during an engine acceleration by the EEC energising the sole-noid control valve vents the P3 servo air from the opening chamber of thebleed valve, and the bleed valve will move to the closed position.

Operating ScheduleThe schedule for one bleed valve − 7C − is shown, in detail, below.

Steady StateIt can be seen that the valve will be commanded closed at stabilised min idle,8600 N2, and will not be opened again in Steady state.

TransientThe valve will be commanded open during engine acceleration whenever N2 isbelow the transient closing speed. Thus during an acceleration from min ” idleto max ” speed the valve will be opened and will remain open until the speedpasses the transient closing speed.If the acceleration is to a speed below the transient closing speed the valve willremain open until the acceleration timer expires ( 30 seconds ).During decelerations the valve will be commanded open whenever N2 is belowthe transient opening speed. The valve remains open until the decelerationceases and a deceleration time, 2 seconds, expires.Note : The transient regime is slightly modified for operation above 15000 ft

but operates in the same way.

Surge / ReverseIf the engine is operating in reverse thrust operation is the same as Transient but different speeds apply. In the event of an engine surge the valve will becommanded open, if the speed is below the open speed, and will remain openuntil the engine restabilises.During an engine deceleration the reverse operation occurs and the bleedvalve opens.

BLEED VALVE OPERATING SCEDULE

BLEED VALVE REGIME OPEN (RPM) CLOSE (RPM)

7A STEADY STATE 11400 11800(35000FT & BELOW)

11800 12250(42000FT & ABOVE)

SURGE & REVERSE

12562 12772

7B STEADY STATE 7650 8000

7C STEADY STATE 6800 7000

TRANSIENT 11600 12050

SURGE &REVERSE

12352 12562

10 STEADY STATE 7650 8000

SURGE &REVERSE

10667( OPEN BELOW )

10667

IDLE 57%=8800 RPM MAX. TO 100% =14950 RPM

Handling bleed valves ( surge bleed )The bleed valves and the solenoid control valves all operate in the same man-ner.FAIL SAFE POSITION : ” 7th and 10th OPEN ”.

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NOTE: FOR TROUBLE SHOOTING PUR-POSES A PNEUMATIC TESTSET ISAVAILABLE TO TEST THE OPERA-TION OF THE BLEED VALVES,BE-CAUSE ONLY THE SOLENOIDVALVES ARE MONITORED !

Page: 313Figure 157 HBV OPEN/CLOSED Schematic

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HANDLING BLEED VALVE MALFUNCTIONSA engineering order ( 010169 ) is released to cover this problems.

7TH / 10TH STAGE HANDLING BLEED VALVES STICKINGHung starts or starting stalls experienced due to 7th and 10th stage handlingbleed valves failing to open or close.The consequences of the malfunction of one or more handling bleed valve‘s on : the ground and airstart capability, the engine operability ( surge free operation ) the engine performance ( EGT, fuel consumption )

have been assessed and are summarized in the following tables :

NOTE: A BLEED TEST SET IS PROVIDED TO CHECK THE BLEEDVALVES AND SOLENOID VALVES FOR PROPER FUNCTION.

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AIRGENERAL

A319/A320/A321IAE V2530−A5

75−00


CONTROL SOLENOID LOCATION

A/C HIGH STAGE BLEED VALVE SOLENOID

7A 7C 7B

HANDLING BLEED VALVE SOLENOIDS

HANDLING BLEED VALVE SOLENOID

10

MAKE UP AIRVALVESOLENOID

Page: 315Figure 158 Bleed Control Valve Solenoids

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BLEED VALVE LOCATIONSThe bleed valves are arranged radially around the HP compressor case asshown below.

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COSTOMER BLEED

SEAL

SILENCER

7A BLEED VALVE

7B (LOWER) BLEED VALVESTAGE 10 BLEED VALVE

7C BLEED VALVE

P3 PRESS CONNECTION

P3 PRESS CONNECTION

Page: 317Figure 159 Bleed Valve Locations

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AIRGENERAL

A319/A320/A321IAE V2530−A5

75−00


HANDLING BLEED VALVE MALFUNCTIONS

A engineering order ( 010169 ) is released to cover this problems.

7TH / 10TH STAGE HANDLING BLEED VALVES STICKINGHung starts or starting stalls experienced due to 7th and 10th stage handlingbleed valves failing to open or close.The consequences of the malfunction of one or more handling bleed valve‘s on : the ground and airstart capability, the engine operability ( surge free operation ) the engine performance ( EGT, fuel consumption )

have been assessed and are summarized in the following tables :

NOTE: A BLEED TEST SET IS PROVIDED TO CHECK THE BLEEDVALVES AND SOLENOID VALVES FOR PROPER FUNCTION.

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Page: 319FRA US/T Bu August 2001 Page: 319Figure 160 HDLG Bleed Valves Malfunction Tables

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Page: 320FRA US/T Bu August 2001 Page: 320Figure 161 Bleed Valve Functional Test

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Page: 321FRA US/T Bu August 2001 Page: 321Figure 162 Bleed Valve Functional Test(cont)

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NACELLE VENTILATIONVentilation is provided for the fan compartment Zone 1, and the core compart-ment Zone 2 to : prevent accessory and component overheating prevent the accumulation of flammable vapours.

Zone 1 VentilationRam air enters the zone through an inlet located on the upper L.H. side of theair intake cowl.The air circulates through the fan compartment and exits at theexhaust located an the bottom rear centre line of the fan cowl doors.

Zone 2 VentilationThe ventilation of Zone 2 is provided by air exhausting from the active clear-ance control ( A.C.C. ) system around the turbine area.The air circulatesthrough the core compartment and exits through the lower bifurcation of the ” C ” ducts.

Ventilation during Ground RunningDuring ground running local pockets of natural convection exist providing someventilation of the fan case - Zone 2.Zone 2 ventilation is still effected in the same way as when the engine is run-ning.

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Page: 323FRA US/T Bu August 2001 Page: 323Figure 163 Nacelle Ventilation

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ATA 75-41 NACELLE TEMPERATURE

NACELLE TEMPERATURE GENERAL

The Nacelle Temperature Sensor has a Measurement Range of −54��C�to 330 �CThis Signal is fed to the EIU which Transforms the Information to digital Form.The EIU Transmits the Data to the ECAM System.The nacelle temperature is displayed if the system is not in engine startingmode and one of the two temperatures reaches the advisory threshold.A advisory indication will be created on the engine system page when the tem-perature reaches approx. 300 - 320 �C.

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NACELLE TEMPERATURESENSOR

CONNECTOR PLUG

EIU

FWC1

FWC2

DMC1 DMC2 DMC3

0.8

1.2

0.8

1.2

LOWER ECAM

300 300

Page: 325Figure 164 Nacelle Temperature System

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IGNITIONGENERAL

A319/A320/A321IAE V2530−A5

74−00


ATA 74 IGNITION74-00 IGNITION SYSTEM PRESENTATION

General

System Operation Dual ignition is automatically selected for:

− all inflight starts− manual start attempts− continuous ignition

Single alternate ignition is selected for ground auto starts.

System TestThe system can be checked on the ground, with the engine shutdown, throughthe CFDS maintenance menu.

IGNITION SYSTEM COMPONENTS

The system comprises: one ignition relay box two ignition exiter units two igniter plugs - located in the combustion system adjacent to No‘s 7&8

fuel spray nozzles. two air cooled H.T. ignition connector leads (cooling is provided by fan air).

Ignition relay boxThe ignition sytem utilises 115V AC supplied from the AC 115V normal andstandby bus bars to the relay box.The 115V relays which are used to connect / isolate the supplies are located inthe relay box and are controlled by signals from the EEC.

NOTE: THE SAME RELAY BOX ALSO HOUSES THE RELAYS WHICHCONTROL THE 115V AC SUPPLIES FOR P2/T2 PROBE HEATING.

NOTE: ACCORDING TO M.E.L. THE IGNITION SYSTEM „ A“ IS RE-QUIRED AS MINIMUM!

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IGNITIONGENERAL

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74−00


IGNITOR PLUG

HIGH TENSION LEAD IGNITION EXCITER 2 (B)

AIR INLET HOSE

B

COOLING JACKET

IGNITION RELAY BOX

IGNITION EXCITER 1 (A)

CH BCONNECTOR

CONNECTOR

CH A

P2/T2 HEATINGCONNECTOR

IGN A IGN BCONNECTOR

Page: 327Figure 165 Ignition System Components

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IGNITIONGENERAL

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IGNITION / STARTING− OPERATION

DescriptionThe ignition circuit is supplied with 115VAC − 400Hz. The electrical power is supplied via the EEC and EIU which controls the ignition of the igniter plugs.A dormant failure of an ignition exciter is not possible for more than oneflight because:− the two ignition systems are independent− the EEC selects alternately ignition system A or B.

FAIL SAFE POSITION: ” IGN RELAYS ,IGN ON”

Ignition during Automatic Start SequenceWhen an automatic start sequence has been activated by the EEC (ENG/MODE selector switch in IGN/START position and MASTER control switch toON),the EEC energizes automatically the appropriate ignition exciter when N2reaches between 10%-16% depending on TAT and keeps it energized until N2reaches 43%.For inflight restart the EEC selects simultaneously both ignition excitersOn the ground, after engine start, the selector must be placed in NORM position, then back to IGN/START to select continuous ignition.( both ignitors)In flight after engine restart, if the selector is maintained in IGN/START position, the EEC selects the continuous ignition on the correspondingengineIn case of a fault during an automatic starting on the ground, the EECaborts automatically the sequence by closing the starter shut−off valve andthe HP fuel shut−off valve and deenergizing the ignitors.

Ignition during Alternate Start Sequence(Manual Start Procedure)When a manual start sequence has been activated by the EEC (ENG/MODEselector switch in IGN/START position and the ENG/MAN START pushbuttonswitch selected to ON) the EEC energizes both ignition exciters.The deenergization of the ignition exciters is automatically commanded by theEEC when engine N2 speed reaches 43%.( Starter cut-out )Positioning of the MASTER control switch to OFF , during that starting sequence, results in ignition exciter deenergization.

Continuous Ignition SelectionManual SelectionWhen the engines are running on the ground or in flight the continuousignition is obtained by positioning the ENG/MODE selector switch inIGN/START position.

Automatic selectionThe EEC selects automatically the continuous ignition in some specific conditions: engine running and air intake cowl anti−icing is selected to ON EIU failed. take−off or during flexible take off approach idle selected. In flight, when there is an engine flameout or stall Reverse

Igniter Plug TestThe operation of the igniter plugs can be checked on the ground, enginenot running, through the maintenance MENU mode of the FADEC or manually ( Manual Start without air )

IGNITION SYSTEM CIRCUIT BREAKERSThere are 5 ignition CB’s installed in the cockpit. 49VU and 121VU

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EEC

121VU

Page: 329Figure 166 Ignition and Starting System Eng. 1

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IGNITION SYSTEM TEST

Igniter Plug TestThe operation of the igniter plugs can be checked on the ground, enginenot running, through the maintenance MENU mode of the FADEC.The test will be performed by selecting the corresponding IGNITOR TESTpage in the MENU and positioning the MASTER control switch to ON to havethe 115VAC power supply to the relevant engine.

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CONTINIUOUE NEXT PAGE

ON

OFF

MASTER 1

ENG

1

I

Page: 331Figure 167 FADEC Ignition Test

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IGNITIONGENERAL

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IGNITOR TEST

Operational Test of the Ignition System with CFDS

Each ignition system must be individually selected to be tested.For the test procedure, refer to AMM TASK 74−00−00−710−041

NOTE: DURING THE TEST,AN AURAL CHECK OF THEIGNITOR PLUG OPERATION HAS TO BEDONE.

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THE GROUND CREW MUST CONFIRM THATTHE IGNITION OPERATES !

Page: 333Figure 168 FADEC Ignition Test Cont.

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IGNITION SYSTEM TEST

Igniter Plug TestThe operation of the igniter plugs can be checked on the ground, enginenot running, through the maintenance MENU mode of the FADEC.The test will be performed by selecting the corresponding IGNITOR TESTpage in the MENU and positioning the MASTER control switch to ON to havethe 115VAC power supply to the relevant engine.

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CONTINIUOUE NEXT PAGE

ON

OFF

MASTER 1

ENG

1

I

Page: 335Figure 169 FADEC Ignition Test

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IGNITIONGENERAL

A319/A320/A321IAE V2530−A5

74−00


IGNITOR TEST

Operational Test of the Ignition System with CFDS

Each ignition system must be individually selected to be tested.For the test procedure, refer to AMM TASK 74−00−00−710−041

NOTE: DURING THE TEST,AN AURAL CHECK OF THEIGNITOR PLUG OPERATION HAS TO BEDONE.

Page: 336

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IGNITIONGENERAL

A319/A320/A321IAE V2530−A5

74−00


THE GROUND CREW MUST CONFIRM THATTHE IGNITION OPERATES !

Page: 337Figure 170 FADEC Ignition Test Cont.

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IGNITIONGENERAL

A319/A320/A321IAE V2530−A5

74−00


IGNITION TEST WITHOUT CFDS

For the test procedure, refer to AMM TASK74−00−00−710−041−01

During the test,an aural check of the ignitor plug operation has to be done.

WARNING: MAKE SURE THAT THERE IS ZERO PSI AT THESTARTER VALVE INLET BEFORE YOU PUSHTHE MAN START P/B.READ THE PRESSURE ONTHE ECAM START PAGE.

Page: 338

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IGNITIONGENERAL

A319/A320/A321IAE V2530−A5

74−00


ENG1

NORM

2. MODE SELECTOR TO− IGN/START

3. MAN START P/B TO− ON

4. MASTER LEVER− ON

1. CHECK AIR PRESSURE AT START VALVE − 0

115VU

ON

OFF

IGN A & B is ”ON”

Page: 339Figure 171 Ignition Test without CFDS

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


ATA 80 STARTING

GENERALStarting SchematicThe starting system of the engine utilizes pressurized air to drive a turbine athigh speed. This turbine drives the engine high pressure rotor through a reduc-tion gear and the engine accessory drive system.The air which is necessary to drive the starter comes from :

− either the APU− or the second engine− or a ground power unit.

The starter supply is controlled by a starter shut−off valve (SOV) pneumatically operated and electrically controlled. In case of failure, the SOVcan be operated by hand.The starter valve closes when the N2 speed reaches 43 %.The starter centrifugal clutch disengages when N2 speed is higher than 43%.Engine starting is controlled from the ENG start panel 115VU located on centerpedestal and ENG/MAN START switch on the overhead panel.The starting sequence may be interrupted at any time by placing the MASTERcontrol lever in OFF position which overrides the FADEC. When the MASTERcontrol lever is in OFF position the HP fuel shut off valve is closed and theengine is stopped.Two procedures are applicable for engine starting : A. Normal Starting Procedure (automatic) The starting sequence is fully controlled by the FADEC and is selected when the ENG/MODE/CRANK/NORM/IGN START selector switch is in IGN/START position and the MASTER control lever in ON position. Start can be aborted on ground only by the FADEC in case of failure. B. Alternative Starting Procedure This sequence controlled by the pilot is as follows: − the ignition selector switch in IGN/START position and MAN START pushbutton switch command the starter shut−off valve, − the MASTER control lever controls the HP fuel shut−off valve.

NOTE: NO START ABORT BY THE FADEC IN CASE OF FAILURE.

Page: 340

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00

Page: 341FRA US/T Bu August 2001 Page: 341Figure 172 Starting System Schematic

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


STARTING COMPONENTS

Starter MotorThe pneumatic starter motor is mounted on the forward face of the externalgearbox and provides the drive to rotate the H.P. compressor to a speed atwhich light up can occur.Attachment to the gearbox is done by a V−clamp adaptor.The starter motor is connected by ducting to the aircraft pneumatic system.The starter motor gears and bearings are lubricated by an integral lubricationsystem. Servicing features include:− oil level sight glass oil fill plug oil drain plug with magnetic chip detector

Starter Motor - OperationThe starter is a pneumatically driven turbine unit that accelerates the H.P. rotorto the required speed for engine starting. The unit is mounted on the front faceof the external gearbox.The starter, shown below, comprises a single stage turbine, a reduction geartrain, a clutch and an output drive shaft − all housed within a case incorporatingan air inlet and exhaust.Compressed air enters the starter, impinges on the turbine blades to rotate theturbine, and leaves through the air exhaust. The reduction gear train convertsthe high speed, low torque rotation of the turbine to low speed, high torquerotation of the gear train hub. The ratchet teeth of the gear hub engage the pawls of the output drive shaft totransmit drive to the external gearbox, which in turn accelerates the engineH.P. compressor rotor assembly.When the air supply to the starter is cut off, the pawls overrun the gear trainhub ratchet teeth allowing the turbine to coast to a stop while the engine H.P.turbine compressor assembly and, therefore, the external gearbox and starteroutput drive shaft continue to rotate. When the starter output drive shaft rota-tional speed increases above a predetermined r.p.m., centrifugal force over-comes the tension of the clutch leaf springs, allowing the pawls to be pulledclear of the gear hub ratchet teeth to disengage the output drive shaft from theturbine.

Starter Air Control ValveThe starter air control valve is a pneumatically operated, electrically controlledshut−off valve positioned on the lower right hand side of the L.P. compressor(fan) case.The start valve controls the air flow from the starter air duct to the starter mo-tor.The start valve basically comprises a butterfly type valve housed in a cylin-drical valve body with in−line flanged end connectors, an actuator, a solenoidvalve and a pressure controller.A micro switch provides valve position feed back information to the FADEC.

Page: 342

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


STARTER

SIGHT GLASS

DRAIN PLUG/CHIP DETECTOR

FILL PLUG

STARTER VALVE

STARTER DUCT

GEARBOX

Page: 343Figure 173 Starter Motor

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


STARTER AIR CONTROL VALVE

DescriptionThe start air control valve is a pneumatically operated , electrically controlledshut−off valve positioned on the lower right hand side of the L.P. compressor( fan ) case

Manual OperationThe starter air valve can be opened/ closed manually using a 0.375 inchsquare drive. Acces is through a panel in the R. H. fan cowl. A valve positionindicator is provided on the valve body.A micro switch provides valve position feed back information to the FADEC.

NOTE: DO NOT OPERATE THE VALVE MANUALLY WITHOUT POSITIVEDUCT PRESSURE.

FAIL SAFE POSITION: ”SOV CLOSED”

Page: 344

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


OP

CL

AMANUAL OVERRIDE

STARTER VALVE

STARTER VALVE

STARTER VALVEFILTER

Page: 345Figure 174 Starter Air Control Valve

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


STARTER AIR CONTROL VALVE

DescriptionThe start air control valve is a pneumatically operated , electrically controlledshut−off valve positioned on the lower right hand side of the L.P. compressor( fan ) case

Manual OperationThe starter air valve can be opened/ closed manually using a 0.375 inchsquare drive. Acces is through a panel in the R. H. fan cowl. A valve positionindicator is provided on the valve body.A micro switch provides valve position feed back information to the FADEC.

NOTE: DO NOT OPERATE THE VALVE MANUALLY WITHOUT POSITIVEDUCT PRESSURE.

FAIL SAFE POSITION: ”SOV CLOSED”

Page: 346

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


OP

CL

AMANUAL OVERRIDE

STARTER VALVE

STARTER VALVE

STARTER VALVEFILTER

Page: 347Figure 175 Starter Air Control Valve

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


START AIR CONTROL VALVE TEST

Start Air Control Valve Test via CFDSThe start air control valve operation may be tested via CFDS.Refer to AMM Task 80−13−51−710−040.

Page: 348

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


NO

TE

:RE

TU

RN

NO

FAU

LTS

OR

RE

TU

RN

FAU

LT D

ET

EC

TE

D

/FM

U T

ES

T

Page: 349Figure 176 Starter Valve Test via CFDS

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


START AIR CONTROL VALVE TEST ( FAULT DETECTED )

AMM Starter Valve Test ata 80-13-51 p507

Page: 350

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


RE

TU

RN

FAU

LT DE

TE

CT

ED

/FM

U T

ES

T

Page: 351Figure 177 Starter Valve Test via CFDS

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


CRANKING−DESCRIPTION

Air SupplyThe air necessary for the starting comes from the duct connecting enginebleed and the precooler..The air necessary for the starter is supplied by either: the other engine through the crossbleed system the APU and in that case, all the air bled from the APU is used for

starting an external source able to supply a pressure between 30 and 40 psig.

Dry Cranking ( Test No 1 )RequirementA dry motoring of the engine will be needed when:

it is necessary to eliminate any fuel accumulated in the combustion chamber

a leak ckeck of engine systems is needed.To perform this operation, the starter is engaged and the engine is motored butthe HP fuel shut off valve remains closed and both ignition systems are OFF.

An engine dry motoring can be performed for a maximum of three consecutive cycles (2 of 2 minutes and 1 of 1 minute with a cooling period of 15 seconds between each cycles). After three cycles or 4 minutes of continuous cranking, stop for a cooling period of 30 minutes.

Page: 352

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


PULL C/B: HP FUEL SOV

CHECK STARTER AIR PRESSURE

PUT MODE SELECTORTO ‘CRANK‘ POSITION

PUSH ‘MAN START‘ PB TO ‘ON‘

MONITOR INDICATIONS

RELEASE ‘MAN START‘ PB TO OFF

PUT MODE SELECTORTO ‘NORM‘ POSITION

PUSH C/B: HP FUEL SOV

LP FUEL SOV OPENS (ECAM WARNING)

ECAM ENG START PAGE APPEARS

MIN. 30 PSI

START VALVE OPENS

N2 AND N1 COMES INTO VIEW

N2, N1 AND OIL PRESSURE MUSTINCREASE

AFTER MAX. 2 MINUTES

ECAM ENG START PAGEDISAPPEARS

LP FUEL SOV CLOSES

NORM

PUSH ONE L/H BOOST PUMP P/BTO ‘ON‘ BOOST PUMP STARTS TO RUN

START VALVE CLOSES,ENGINE INDICATIONS−BACK TO ‘0‘

Page: 353Figure 178 Dry Cranking Procedure

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


WET CRANKING

Wet Cranking ( Test No 2 )A wet motoring will be needed when the integrity of the fuel system has tobe checked.If such a test is performed, both ignition systems are off ( also pull the circuitbreakers) and the starter is engaged to raise N2 up to the required speed of20%.The MASTER control switch is moved to ON and the exhaust nozzle of the engine carefully monitored to detect any trace of fuel. On the ECAM the FFindication shows approx. 180kg initial fuel flow.When the MASTER control switch will be returned to the OFF position toshut-off the fuel , also the starter valve closes . The EEC automaticallyreengages the starter at 10% N2 and the engine should be motored for atleast 60 seconds to eliminate entrapped fuel or vapor.

The motoring can be performed for a maximum of three consecutivecycles (2 of 2 minutes and 1 of 1 minute with a cooling period of 15seconds between each cycles).After three cycles or 4 miutes of continuous cranking, stop for a cooling period of 30 minutes.

Page: 354

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


CHECK STARTER AIR PRESSURE

PUT MODE SELECTORTO ‘CRANK‘ POSITION

PUSH ‘MAN START‘ PB TO ‘ON‘

MONITOR INDICATIONS

ECAM ENG START PAGE APPEARS

MIN. 30 PSI

START VALVE OPENS

N2 AND N1 COMES INTO VIEW

N2, N1 AND OIL PRESSURE MUSTINCREASE

PUSH ONE L/H BOOST PUMP P/BTO ‘ON‘ BOOST PUMP STARTS TO RUN

WHEN N2 SPEED IS >20%

PUT ENG MASTER SWITCH TO ‘ON‘ FUEL FLOW INDICATION INCREASES

PULL IGNITION SYSTEM C/B‘S (5)

AFTER 10−20 SECONDS

PUT ENG MASTER SWITCH TO ‘OFF‘ FUEL FLOW INDICATION GOES TO ‘0‘START VALVE CLOSES

WHEN N2 SPEED REACHES 10% THE EEC RE−ENGAGES THE STARTER

AFTER 60 SECONDS MOTORING

RELEASE ‘MAN START‘ PB TO OFF

PUT MODE SELECTORTO ‘NORM‘ POSITION

ECAM ENG START PAGEDISAPPEARS

START VALVE CLOSES,ENGINE INDICATIONS−BACK TO ‘0‘

NORM

DO NOT PULL C/B: HP FUEL SOV

Page: 355Figure 179 Wet Cranking Procedure

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


AUTOMATIC STARTThe automatic start mode gives the EEC full control to automatically sequencethe starter air valve, ignition relays and the fuel on / off torque motor. Upon re-ceipt of the appropriate start command signals from the engine interface unit( EIU ) , the EEC commands , in sequence: the starter air valve ignition exiter relay(s),

− alternatively selected for each ground start− both selected for inflight or manual starts

fuel on function of the torque motor which opens the shutoff valve.During a normal start, the starter air valve and ignition exciter are automaticallyturned off by the EEC at a predetermined N2 speed of 43%Starter assist will be comanded by the EEC for inflight starts at low MACHnumbers where windmilling conditions are insufficient for engine starting.(The EEC has input data necessary to activate starter assist function wherenecessary.)

NOTE: IN CASE A AUTO START IS INITIATED AND ONE THRUST LEVERIS NOT IN IDLE POSITION A ECAM WARNING IS TRIGGERT. THESTART SEQUENCE WILL CONTIUE AND THE ENGINE WILL AC-CELERATE TO THE TRUST LEVER POSITION.

EEC AUTO START ABBORTThe autostart procedure commences only when the engine is not running, themode selector set to IGN/START and the master switch is ON.Intermittent mode selector position or manual start push button switch selectionhas no effect on autostart sequence once the autostart procedure is initiated.Switching the master switch OFF during an autostart will close the fuel andstarter air valves and turn the ignition system off.It also resets the EEC.

The automatic start abort function is only available when N2 speed is below43% and in case of: Start valve failure Ignition failure Pressure Raising Shut Off Valve failure Hot start Hung start Surge EGT >250 deg C when restart (max 2 min) Loss of EGT

NOTE: THE OIL PRESSURE IS NOT MONITORED DURING AUTO STARTThe EEC automatically shuts off fuel, ignition, and starter air and provides theappropriate fault indication to the cockpit. (Auto Start Fault)Autostart fault messages will be displayed until approximately idle speed.The EEC’s ability to shut off fuel is inhibited above 43% N2 on theground and at all conditions inflight. In case of an automatic start abort,the EEC re−opens the start valve when reaching 10% N2 for a 30 seconddry motoring cycle to clear fuel vapor and to cool the engine.Then the operator has to select the Master switch to the OFF position bya command indicated on the ECAM page ( ”Master lever OFF ” ).The operator then has to decide to perform a new engine start or trouble-shoot the system.

Page: 356

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


Panel 115 VU

−Turn Mode Selector to IGNSTART Position

ECAM ENG Start Page is displayed, the airpreessure( HP−Connection or APU ) must be 30−40 psi.

ENG1

Panel 115 VU

−Set the ENG−MASTER switch to ON On the ENG Start Page: − the starter valve symbole goes in line (open)After 30 seconds:-the A or B IGN indication comes in to view-the FUEL FLOW indication 180KG/H comes intoview -the EGT rises (max. 20 sec. after FF).

NORM

NORM

Upper ECAM−MONITOR: EPR, N1, N2, EGT, FF

−at 43% N2 the starter valve symbole must go to cross line (closed)− IGN OFF−Check Oil Pressure min. 60psi.−record the start EGT (R/U sheet)

Panel 115 VU

−Turn Mode Selector to NORM

NORM

ENG1

( The Pack valves also ”Close” )

Page: 357Figure 180 Automatic Start Procedure

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


MANUAL STARTThe engine manual start panel, used for manual start, is located on the over-head panel and is composed of two manual start push button switches (one perengine).The manual start mode limits the authority of the EEC so that the pilot can sequence the starter, ignition and fuel on/off manually. This includes the abilityto dry crank or wet crank.During manual Start operation, the EEC Auto Startabort feature is notavailable and conventional monitoring of the start parameters is required.The EEC continues to provide fault indications to the cockpit.

The manual start procedure commences when the mode selector is set to:IGN/START,the manual start push button switch is set to ON and the masterswitch is OFF. The starter air valve is then commanded open by the EEC.When the master switch is turned ON ( at 22% N2 ) during a manual start, bothignitors are energized ( IGN A/B ) and fuel is turned on ( Intial FF 180 KG/H). Intermittent mode selector position has no effect on the manual start sequenceonce the manual start procedure is initiated.The starter air valve can be closed by selecting the manual start push buttonswitch OFF at any time prior to turning the master switch ON.Once the master switch is turned ON, the manual start push button switch hasno effect on the start.When the master switch is turned OFF, the control commands the HP fuelvalve closed, the starter air valve closed and the ignitors off and the EEC isresetted..

Page: 358

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STARTINGGENERAL

A319/A320/A321IAE V2530−A5

80−00


NORM

Panel 115 VU Panel 122 VU − Turn MODE Selector to IGN / START ECAM ENG START Page is displayed, the airpressure − Turn MODE Selector to NORM

Position ( HP - Connection or APU ) must be 30 - 40 psi

Panel 122 VU−Push the MAN START PB − the blue ON light of this PB comes on.

On the ENG Start Page : −the starter valve symbole goes in line (open).

−N2, Oilpressure and N1 must increase

NORM

Panel 115 VU

−after 30sec (> 22% N2):

set the ENG MASTER switch to ON

ENG1

−A and B indication comes in to view ( below IGN )

−FUEL FLOW indication 180KG/H

−EGT rise (max. 20 sec. after FF )

−at43% N2 the starter valve symbole must go to cross line (closed)− IGN OFF−Check Oil Pressure min. 60psi.−record the start EGT (R/U sheet)

NORM

ENG1

Panel 122 VU

−release the MAN START PB(−PACK VALVES closed )

Page: 359Figure 181 Manual Start Procedure

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ExhaustReverser System

A319/A320/A321IAE V2530-A5

78−30


ATA 78 EXHAUST

REVERSER SYSTEM

Introduction

DescriptionThe thrust reverser comprises a fixed inner and a movable outer ( translating )assembly.The translating cowl is moved by four hydraulically operated actuators whichare pressurized by the pumps mounted on each engine..The air is discharged through cacades.The reverser is controlled through the FADEC system from the cockpit by alever hinged to the corresponding throttle control lever-The thrust reverser system comprises:

− a hydraulic control unit (HCU)− four actuators with internal lock for lower actuators− three flexible shafts− two linear variable differential transformers located on each upper actuator− two proximity switches located on each lower actuator− two thrust reverser cowls comprising a fixed structure and 2 translating sleeves latched together.

Page: 360

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A319/A320/A321IAE V2530-A5

78−30


DRAG LINK

Page: 361Figure 182 Thrust Reverser stowed / deployed

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A319/A320/A321IAE V2530-A5

78−30


THRUST REVERSER SYSTEM DESCRIPTION

GeneralThe thrust reverser is actuated in response to signals from the Engine Electronic Control (EEC). Selection of either stow or deploy from thecockpit generates a signal to the engine EEC which in turn, suppliessignals to the thrust reverser hydraulic control unit.

Thrust Reverser DeploymentThrust reverser deployment is initiated by rearward movement of thereverser lever which inputs a signal, via a dual resolver, to the EEC.The EEC supplies a 28 volt signal to the isolation valve and directional control valve solenoids mounted in the HCU.The supply of the signal to the directional control valve solenoid is also depen-dent if aircraft is on ground (weight onwheels) and upon the closure of the air-craft permission switch ( T/R inhibition relay) in that line. This switch is closedby the Throttle Lever Angle signal via the spoiler/elevator computer and theEngine Interface Unit energization of the isolation valve solenoid and the direc-tional control valve solenoid allows hydraulic pressure into the system .Thisevent being relayed to the EEC by the pressure switch mounted in the HCU.Pressure in the lower actuators releases the locks and these events are sig-nalled to the EEC by the Proximity Switches (lock sensors). As the pistonsmove rearward to deploy the reverser, the Linear Variable Differential Trans-former (LVDT) on the upper actuators monitors the movement and informs theEEC when the translating sleeve is fully deployed, the Proximity Switches andLVDTs remain active and the isolation valve remains energized.

Thrust Reverser StowageStowage of reverser is initiated by forward movement of the piggybacklevers which signal this intent to the EEC. The signal to the directionalcontrol valve solenoid is then cancelled by the EEC and permission switch,allowing pressure to remain only in the stow side of the actuators. The pistonsthen move forward until stowing is complete and the lower actuator locks areengaged after which the isolation valve solenoid is de−energized and the re-verser is locked in the forward thrust mode.

NOTE: DURING NORMAL REVERSER OPERATION THE ISOLATIONVALVE REMAINS ENERGIZED FOR A PERIOD OF FIVE SE-CONDS AFTER THE LVDTS HAVE REGISTERED FULLY STOWEDTO ENSURE FULL LOCK ENGAGEMENT AND COMPLETION OFTHE STOW CYCLE.

Inadvertent Stowage/DeploymentIn either case the LVDT sensors would detect a movement the EECwould execute auto−restow or auto−redeploy.This occurs when the LVDTs sense uncommanded movement greater than10% of actuator full travel.When auto−restow is initiated the EEC signals the isolation valve to open. Pressure is returned to the system and with the directional control valvein its stow position the reverser is returned to its stowed condition.Following auto−restow the isolation valve would remain energized for theremainder of the flight.If the reverser travel exceeds 15% of its travel from the fully stowed positionthen the EEC will command idle.Following restow, full power is again obtainable.When auto redeploy is initiated to counteract inadvertent stow, the EEC willcommand the isolation valve to close and maintain it closed until forward thrusthas been reselected. This action will prevent further movement in the stowdirection by virtue of the large aerodynamic loads on the translating sleeveswhich will normally be sufficient to deploy the reverser. If the reverser travelexceeds 22% of its travel from the fully deployed position then the EEC willcommand idle power.

T/R components monitored by CFDSThe following components are monitored by the CFDS: HYDRAULIC CONTROL UNIT (HCU) STOW SWITCH LOWER ACTUATOR R/H STOW SWITCH - LOWER ACTUATOR L/H LVDT -THRUST REV UPPER ACTUATOR R/H ( DEPLOY ) LVDT - THRUST REV UPPER ACTUATOR L/H ( DEPLOY )

Page: 362

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A319/A320/A321IAE V2530-A5

78−30


CFDIU

E.E.C.

TLARESOLVERS

POTENTIO−METERS

THRUSTLEVER

INHIBITION RELAY

EIU 1 2/

MCDUT R TEST

FWC

DMC

CHANNEL A

CHANNEL B

PRESS SW

HYDRAULICRETURN

CHANNEL BCHANNEL A

HCU T R/

/

DIRECT V SOLISOLATION V SOLDIRECT V SOL

ORSEC 1

SEC 2 (3 )

LGCIU 1/2

(WOW)

AND

T/RPOSITION

CHANNEL A

T/RPOSITION

CHANNEL B

N2 >50%

LANDINGGEARS1&2

MAIN

SUPPLY

STATICRELAY

FSOV

REVREV

EPR

1,009 1,010

1

1,2 1,41,6

1

1,2 1,41,6

ISOLATION V SOL

MREV 70,0%

Page: 363Figure 183 Reverser System Schematic

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THRUST REVERSER INDEPENDENT LOCKING SYSTEM

General

**ON A/C 116−199,An independent locking system is designed to isolate the thrust reverser fromthe aircraft hydraulic system. This system consists of thrust reverser Shut−OffValve (SOV) upstream of the Hydraulic Control Unit (HCU), a filter andassociated plumbing, mounting and electrical supply. The SOV is electricallyactuated from an independent signal from the SEC (Spoiler Elevator Com-puter), bypassing the FADEC command circuit.

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CFDIU

E.E.C.

TLARESOLVERS

POTENTIO−METERS

THRUSTLEVER

INHIBITION RELAY

EIU 1 2/

MCDUT R TEST

FWC

DMC

CHANNEL A

CHANNEL B

PRESS SW

HYDRAULICRETURN

CHANNEL BCHANNEL A

HCU T R/

/

DIRECT V SOLISOLATION V SOLDIRECT V SOL

ORSEC 1

SEC 2 (3 )

LGCIU 1/2

(WOW)

AND

T/RPOSITION

CHANNEL A

T/RPOSITION

CHANNEL B

N2 >50%

LANDINGGEARS1&2

MAIN

SUPPLY

STATICRELAY

FSOV

REVREV

EPR

1,009 1,010

1

1,2 1,41,6

1

1,2 1,41,6

ISOLATION V SOL

Page: 365Figure 184 Reverser System Schematic

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THRUST REVERSER SYSTEM

CacadesThere are 16 cacades installed.The cacades are not interchangeable.

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Page: Page: 367FRA US/T Bu August 2001 Page: 367Figure 185 Reverser Installation

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THRUST REVERSER HYDRAULIC SUPPLY

Thrust Reverser OperationThe thrust reverser is operated by aicraft hydraulic pressure.The reverser hydraulic control unit ( HCU ) directs hydraulic pressure to theactuators.The EEC controls the HCU and the reverser operation.

THRUST REVERSER MANUAL DEPLOYMENT

Non Return Valve ( By−pass ).During manual deployment the non return valve must be set in the bypass posi-tion to allow the hydraulic from the actuators to go back to return.Access to the non return valve is gained by removing the pylon access panelon the left hand side..

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NON RETURN VALVE

Page: 369Figure 186 Reverser Hydraulic Supply

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THRUST REVERSER INDEPENDENT LOCKING SYSTEM

**ON A/C 116−199,

GeneralAn independent locking system is designed to isolate the thrust reverser fromthe aircraft hydraulic system. This system consists of thrust reverser Shut−OffValve (SOV) upstream of the Hydraulic Control Unit (HCU), a filter andassociated plumbing, mounting and electrical supply. The SOV is electricallyactuated from an independent signal from the SEC (Spoiler Elevator Com-puter), bypassing the FADEC command circuit.

Component Location The SOV and the filter are located under the pylon. (Ref. Fig. 001)

COMPONENT DESCRIPTION

Shut−Off ValveThe thrust reverser Shut−Off Valve (SOV) is a 3 port, two position spool valve.It is controlled by a solenoid driven 3 port, two position normally open pilotvalve. Electrical power is supplied to the SOV through the fan electrical feederbox.

Filter and Clogging IndicatorIt is used to filter the fluid from the aircraft hydraulic system. The filter is a flow−through cartridge−type filter. The clogging indicatormonitors the pressure lossthrough the filter cartridge and has a pop−out indicator to signal when it is nec-essary to replace the filter element. Two spring−loaded magnetic pistons keepthe pop out indicator in retracted position. The lower magnetic piston monitorsthe differential between the filtered and unfiltered fluid pressure across the filterelement. As the differential pressure increases, the piston compresses itsspring and moves away from the upper magnetic piston. At a preset displace-ment of approximately 2 mm, the upper magnetic piston spring overcomes themagnetic force and drives the pop−out indicator from its retracted position.Thefilter assembly contains a check valve to permit the removal of the canister andthe change of the filter element with a minimum of spillage. LOCATION

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Page: Page: 371FRA US/T Bu August 2001 Page: 371Figure 187 T/R Independent Locking System (**On A/C 116−199)

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REVERSER HYDRAULIC CONTROL UNIT

Reverser Hydraulic Control Unit ( HCU )GeneralThe hydraulic control unit controls hydraulic fluid flow to the thrust reverser ac-tuators. Control and feedback signals are exchanged with the EEC.The HCU is mounted on the pylon over the engine centerline, justforward of the C−duct and is accessible from the left side.The hydraulic control unit includes the following items:

− isolation solenoid valve solenoid,− isolation valve,− directional control valve solenoid,− directional control valve,− pressure switch,− filter and clogging indicator (pop out).

Isolation ValveThe solenoid operated isolation valve isolates the thrust reverser actuation sys-tems from the remaining hydraulic network on the engine. The isolation valvesolenoid is a dual coil valve solenoid connected to both channels of the EEC.The isolationvalve is in the closed position while the thrust reverser is in thestowed position. Upon actuation of the thrust reverser system, the isolationvalve solenoid is energized and the isolation valve is opened.

Directional Control ValveThe solenoid operated directional control valve directs high pressure hydraulicfluid to the correct end(s) of the actuators to either stow or deploy the translat-ing sleeve.The directional control valve solenoid is a dual wound solenoid con-nected to both channels of the EEC. The directional control valve solenoid isenergized when the deploy command is given and provides hydraulic fluid athydraulic pump supply pressure to both ends of the actuators through thedirectional control valve to initiate deployment of translating sleeve.

Pressure SwitchThe pressure switch provides signals to the EEC to indicate when there is hy-draulic pressure downstream of the isolation valve. The pressure switch isclosed at pressure between 798 and 1450 psi and is opened at a minimum-pressure of 798 psi.

Filter and clogging indicatorThe hydraulic control unit filter is used to filter the fluid supply from the aircrafthydraulic system. The filter is a flow through cartridge type filter. The cloggingindicator monitors pressure loss through the filter cartridge and features a pop−out indicator to signal when it is necessary to replace the filter element.

Manual Lockout LeverWith the manual lockout lever it is possible to shut the hydraulic supply to thereverser by closing the isolation valve in the HCU.The lever can be secured inthe lockout position with a pin.(this is also a part of blocking the reverser.)This must always be done when working on the reverser system !

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78−30


FILTER

PRESSURE

SWITCH

HYDRAULIK

OUTPUT

CONNECTOR „A“ CHANNEL

CONNECTOR „B“ CHANNEL

HOUSING

FILTER (POP-OUT)

INDICATOR

SPRING

ISOLATION VALVE

SOLENOID VALVE

DIRECTIONAL VALVE

SOLENOID VALVE

BLEED VALVE

QUICK RELEASE PIN

ISOLATING LEVER

HYDRAULICINPUT UNION

Page: 373Figure 188 Hydraulic Control Unit ( HCU )

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78−30


HCU IN FORWARD THRUST POSITIONIn the initial stowed position with the reverse stow control selected in the cock-pit,the hydraulic pressure is applied to the input of the HCU.All reverser hydrau-lic systems are pressurized at the return pressure as long as the aircraft is inflight and no signal is sent to open the isolation valve solenoid.

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78−30


FORWARD THRUST CONFIGURATIONACTUATORS STOWED.

NON RETURN VALVE ( MANUAL OPERATED)

S

FILTER

**ON A/C 116−199,SHUT-OFF VALVE

Page: 375Figure 189 HCU Schematic

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HCU DEPLOY SEQUENCE DESCRIPTION

1 When reverse thrust is selected in the cockpit, the EEC ensures that deploy ment is permitted.In that case, the electrical power (28VDC) is sent to the isolation valve solenoid and to the directional valve solenoid.

2 When the isolation valve is opened and the directional control valve solenoid is energized, hydraulic pressure (3000 psi) moves the directional control valve to supply hydraulic pressure to the head end of the actuator to unlock the actuators, and then extending the actuators.

3 As soon as both lock sensors indicate unlocked for more than 0.2 seconds (indicating that translating sleeves are ”unlocked sleeves” signal is sent by these sensors to the EEC. In the cockpit an amber REV indication is dis played in the middle of the EPR dial or the ECAM display unit.

4 Each translating sleeve arriving at 95 percent of its travel is slowed down until completely deployed through hydraulic actuator inner restriction. This event is indicated to EEC when both Linear variable Differential Transform ers (LVTD) detect this position. REV indication changes to green.

NOTE: WHEN THE THRUST REVERSER IS IN THE DEPLOYED POSI-TION, THE ISOLATION VALVE REMAINS ENERGIZED TO MAIN-TAIN THE HYDRAULIC PRESSURE IN THE ACTUATORS TO PRE-VENT VIBRATION. IF AN UNCOMMANDED STOW MOVEMENT ISDETECTED, THE EEC WILL DE−ENERGIZE THE ISOLATIONVALVE. THIS WILL LEAD TO A THRUST REVERSER REDEPLOYDUE TO AERODYNAMICAL FORCES ON THE BLOCKER DOORS.

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Page: Page: 377FRA US/T Bu August 2001 Page: 377Figure 190 HCU Deploy Sequence

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HCU STOW SEQUENCE DESCRIPTION

1 When translating sleeves stowing is selected, the EEC ensures that stowing is permitted. In that case the EEC de−energizes the directional valve sole noid. When one translating sleeve is less than 95 % deployed, REV indica tion changes to amber.

2 Hydraulic pressure is supplied to the rod end of the actuator, the head is connected to return. A flow limiter controls hydraulic actuator piston retrac tion speed.

3 When both translating sleeves are at 0 % from their stowed position, they set the proximity switches (lock sensor) which send the ”stowed sleeves” information to the EEC. The REV indication disappears.

4 The actuators move until stowing is complete and the lower actuator locks are engaged after which the isolation valve solenoid is de−energized and the reverser is locked in the forward thrust mode position.

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Page: Page: 379FRA US/T Bu August 2001 Page: 379Figure 191 HCU Stow Sequence

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HYDRAULIC ACTUATION SYS. COMP.

Hydraulic ActuatorsThe actuator base is attached to a torque ring and the end of the piston is at-tached to the translating sleeve. As hydraulic pressure builds up in the actua-tor, the piston extends. This moves the translating sleeve aft to the deploy posi-tion. In the retract mode,the piston retracts which moves the translating back tothe stow position.The Upper actuators ( 2 ) have internal LVDT.The Lower actuators ( 2 ) have a manual unlocking handle and proximityswitches.

FLEXSHAFT INSTALLATION

Syncronization System

Flexible ShaftsThree flexible shafts connect the four actuators together to synchronizethe speed with which the actuators operate and the T/R sleves on each side ofthe engine .This synchronization keeps the top and bottom of the sleeve traveling at thesame rate so the sleeve will not tilt and jam. The synchronization also keepsthe two translating sleeves moving together so reverse pressure in the second-ary air flow is equal on both sides of the engine.The flexible shafts are installed inside the extend (deploy) hydraulic hoses. Theshaft engages a worm gear at the base of the actuator that translates the turn-ing action of the actuator piston as it moves out or in.A cross−over shaft connects the two upper actuators. Another shaft connects the upper and lower actuators on each side.

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MANUAL DRIVE

Page: 381Figure 192 Flexible Drive Shafts

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HYDRAULIC ACTUATORS DESCRIPTIONFour actuators are used for each thrust reverser,two actuators are used foreach translating cowl. the lower actuators incorporate an integral lock mechanism which holds the

piston in the fully stowed position. the upper actuators incorporate an integral Linear Variable Directional

Transformer (LVDT) to indicate piston position,and thus translating cowlposition , to the EEC.

All actuators use hydraulic snubbing at the end of the deploy stroke to slowdown the actuators over the final part of the deploy stroke.All actuators alsoincorporate the necessary deploy stroke mechanical stops.

UPPER NONLOCKING ACTUATORThe two upper actuators are identical and in conjunction with the two lowerlocking actuators , control movement of the fan reverser translating elements inresponse to hydraulic inputs from the HCU.

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Page: Page: 383FRA US/T Bu August 2001 Page: 383Figure 193 Upper Nonlocking Actuator

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LOWER LOCKING ACTUATORSThe two lower loking actuators are identical and in conjunction with the two up-per actuators, control movement of the fan reverser translating elements in re-sponse to hydraulic inputs from the hydraulic control unit (HCU).

The actuators incorporate an integral lock mechanism to hold the piston rodwhen the actuator is in the fully stowed position.The lock releases on rising hydraulic pressure when deploy is commanded viathe HCU.The lock mechanism incorporates a manual release facility and prox-imity switch for electrical lock position feedback to the EEC.

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Page: Page: 385FRA US/T Bu August 2001 Page: 385Figure 194 Lower Locking Actuator

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THRUST REVERSER MANUAL DEPLOY / STOW

Manual Deploy/stowThe thrust reverser may be deployed/stowed manually for maintenance − trou-bleshooting operations.

The procedure is summarised below, the full procedure, warnings and cau-tions may be found in the MM ATA 78−30. open and tag the CB’s listed in the MM. open the L. and R. hand fan cowls. move the thrust reverser hydraulic control unit de−activation lever to the de−

activated position and insert the lockout pin. disengage the locks on the two locking actuators. Insert pins to ensure locks

remain disengaged. position the non return valve in the bypass position ( deploy only−not neces-

sary for stow operation ). insert 3/8 inch square drive speed brace into external socket, push to en-

gage drive and rotate speed brace to extend/retract translating cowl as re-quired.

NOTE: DO NOT EXCEED MAX. INDICATED TORQUE LOADING.

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NON RETURN VALVE

Page: 387Figure 195 Reverser Manual Operation

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78−30


THRUST REVERSER DEACTIVATION

De−activationThe procedure is summarised below, the full procedure is described in the MM 78−30−00 P.407. if the thrust reverser is deployed,it has to be stowed manually. install the lock out pin in the de−activation lever of the hydraulic control unit. remove the translating cowl de−activation pins (2) from their stowage and

insert them in the de−activation position.

T / R Lockout pin installation

NOTE: WHEN FULLY INSERTED IN THE DE−ACTIVATION POSITION THEPINS WILL PROTUDE APPROX. 0.8” TO PROVIDE VISUAL IN-DICATION OF ”LOCK OUT”.

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Page: Page: 389FRA US/T Bu August 2001 Page: 389Figure 196 T/R Deactivation

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FADEC CFDS REVERSER TEST

Reverser Testing via MCDUVia MCDU it is possible to operate the reverser on ground with engines OFFtomake sure the system operation is o.k.For the TEST refer to:MM Task 78−31−00−710−41 Operational Test of the Thrust Reverser Systemwth the CFDS.

DescriptionFor the test hydraulic power must be switched on depending which reversersystem will be tested.( Green ENG 1, Yellow END 2 ).All the test steps are written on the MCDU.If the test is active the REV UN-STOW warning appears on the engine warning display.Movement of the throttle into the reverse idle position will deploy the rever-ser.Returning the throttle to the FWD idle position will restow the reverser.During the test also the REV indication in the EPR indicator must be checked.The actual position of the T/R is also indicated on the MCDU .

CAUTION: MAKE SURE THE TRAVEL RANGES OF THETHRUST REVERSERS ARE CLEAR.FOR SAFTEY REASONS THE TEST TIME DURA-TION IS LIMITED TO 60 SEC.

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WARNING MASSAGE WHEN IN TEST

Page: 391Figure 197 FADEC T/R Test (NO FAULT)

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FADEC T/R TEST ( FAULT DETECTED )

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Page: Page: 393FRA US/T Bu August 2001 Page: 393Figure 198 FADEC T/R Test (FAULT DETECTED)

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FADEC T/R TEST ( NOT O.K. )For saftey reasons the time for the test is limited.

NOTE: IF THE TEST PROCEDURE IS NOT PER-FORMED WITHIN 15 SECONDS (MOVING THETHROTTLE LEVER TO REVERSE ) THE TESTWILL BE INTERRUPTED AND A NEW TESTMUST BE INITIATED.

NOTE: THE DURATION OF THE COMPLETE T/R OP-ERATIONAL TEST (OPENING & CLOSING ) ISLIMITED TO 60 SECONDS. IF THIS TIME IS EXCEEDED THE TEST WILLBE INTERRUPTED AND A NEW TEST MUSTBE INITIATED.

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NO THRUST LEVER MOVEMENTTO REV. WITHIN THE TIMELIMIT

Page: 395Figure 199 FADEC T/R Test (NOT O.K.)

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ENGINEENGINE CHANGE

A319/A320/A321IAE V2530-A5

71-00


ATA 71-00 ENGINE CHANGEENGINE REMOVAL / INSTALLATION

The arrangements for slinging / hoisting the engine are shown below( Bootstrap).

NOTE: DURING THIS OPERATION THE ”C” DUCTS ARE SUPPORTED BYRODS WHICH ARE POSITIONED BETWEEN THE ”C” DUCT ANDTHE ENGINE PYLON.

After a new engine was installed different Test Tasks have to be performed: Check of engine datas via CFDS ( ESN,EEC P/N, Engine Rating, Bump

level ) to make sure that they are the same as written on the EEC, dataentry plug and engine identification plates.

Operational Test of EEC via CFDS. If A/C is operated in actual CAT III conditions,a Land Test must be per-

formed. Functional check of IDG disconnect system. Functional check of engine ice protection system. TEST NO. 1 ( Dry motor leak check ) TEST NO. 2 ( Wet motor leak check ) TEST NO. 3 ( Idle leak check ) TEST NO. 6 ( EEC system idle test ) TEST NO. 13 ( Prestested engine replacement test )

For further information refer to AMM ATA 71-00-00.

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ENGINEENGINE CHANGE

A319/A320/A321IAE V2530-A5

71-00

Page: 397FRA US/T Bu August 2001 Page: 397Figure 200 Engine Removal / Installation

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ENGINE GROUND OPERATIONPOWER PLANT PRESERVATION

A319/320/321

71−00

FRA US/T5 KoA May 04 Page: 398

POWER PLANT PRESERVATION

List of Preservation ProceduresTask 71−00−00−500−010

Preservation of the Power PlantTask 71−00−00−550−010A.General The procedures in AMM 71−00−00−550−010 are for Engines stored on wing or inside/outside.The periods are for 7 days,7−30 days,31 days-3 months and over 3 months.B.Equipment Engine Covers,desiccant,rust preventative,moisture resistant tape.

Depreservation of the Power PlantTask 71−00−00−550−011A.General The procedures in AMM 71−00−00−550−011 are for depreservation up to 3 months or over 3 months.B.Equipment Engine Oil

Clean and Examine the Power PlantTask 71−00−00−100−010A.General The procedures in AMM 71−00−00−100−010 describes the cleaning of the Engine.B.Equipment Cleaning Solvent

Protect the Engine External SurfacesTask 71−00−00−600−011A.General The procedures in AMM 71−00−00−600−011 describes the protection of external surfaces.B.Equipment Anti-corrosion inhibit fluid.

Preservation of the Main Line BearingsTask 71−00−00−550−012A.General The procedures in AMM 71−00−00−550−012 describes the preservation of the main line Bearings.B.Equipment Engine Oil.

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ENGINE GROUND OPERATIONPOWER PLANT PRESERVATION

A319/320/321

71−00

FRA US/T5 KoA May 04 Page: 399

THIS PAGE INTENTIONALLY LEFT BLANK

Page: 399

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ICE AND RAIN PROTECTIONENG. AIR INTAKE ANTI ICE PROTECTION

A319/A320/A321IAE V2530−A5

30−20


ATA 30 ICE AND RAIN PROTECTION

30−20 ENG. AIR INTAKE ICE PROTETION

System Description

Engine Air Intake Anti−Ice Air sourceThe air bled from the 7th stage of the high compressor is the heat source.A solenoid−operated shutoff valve (which is designed to fail to the open position) provides the on−off control. The piccolo tube distributes the air whithin the leading edge of the intake cowl. The spent air exhausts via aflush duct in the aft cavity of the intake cowl.

ValveFor each Engine, hot bleed air is ducted via an ”ON/OFF” valve.The valve is pneumatically operated,electrically controlled and spring loadedclosed.Upon energization of the solenoid, the valve will close.In case of loss of electrical power supply and pneumatic air supply available,the valve will open.(Fail safe „OPEN“ ) It has a “Manual Override and Lock”. It can be blocked in the OPEN or in

the CLOSED position.

ControlFor each engine, the”ON/OFF” valve is controlled by a pushbutton.Continuos ignition (A/B) is automaticaly activated on both engines when thevalve is opened.The ”FAULT” light comes on during transit or in case of abnormal operation.When the anti−ice valve is open, the zone controller determines the bleed airdemand for the Full Authority Digital Engine Control (FADEC) system.

ECAM PageIf at least one of the two engine air intake anti−ice systems is selected ”ON”, amessage appears in GREEN on the ”ECAM MEMO” display.

SYSTEM CONTROL

ON − (PB−Switch In, Blue)The ON light comes on in blue. (valve solenoid deenergized) .ENG ANTI ICE ON is indicated on the ECAM MEMO page.When the anti ice valve is open (valve position sw. NOT CLOSED), the zonecontroller sends a signal to the FADEC (ECS signal), this will: Modulate the Idle speed to Min.PS3 Schedule Demand for both engines. Switch the Cont. Ignition− ON (via EIU/EEC).

OFF − (PB−Switch Out) Anti ice system is OFF (valve solenoid energized).

FAULT − (PB Switch In, Amber)Fault light illuminates amber when valve not fully open.

FAULT − (PB−Switch Out, Amber)Fault light illuminates amber.The ECAM is activated − Single chime sounds − MASTER CAUT light ”ON” − Warning message:

− ANTI ICE ENG 1 (2) VALVE CLSD− ANTI ICE ENG 1 (2) VALVE OPEN.

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A319/A320/A321IAE V2530−A5

30−20


1 2

7

OPEN POSITION SIGNAL CABIN ZONE

CONTROLLER

FADEC

Page: 401Figure 201 Engine Nacelle A/I Architecture

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A319/A320/A321IAE V2530−A5

30−20


SYSTEM CONTROL SCHEMATIC

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A319/A320/A321IAE V2530−A5

30−20


( EIU )

( ZONE CONT.)

Page: 403Figure 202 Control Schematic

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A319/A320/A321IAE V2530−A5

30−20


ENGINE ANTI ICE DUCT AND VALVEANTI−ICE VALVE DEACTIVATION

refer to MEL.ATA 30.

Procedure Lock the intake anti−ice valve (1) in the open or the closed position Remove the lock−pin (4) from the transportation hole (5) in the

valve (1). Use an applicable wrench on the nut (2) and move the valve to the

necessary position (open or closed). Hold the valve in the necessary position and install the lock−pin

(4) in to the valve locking hole (3).

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A319/A320/A321IAE V2530−A5

30−20


ANTI−ICE DUCT

ANTI−ICE VALVE

5 TRANSPORTATION−

HOLE

4 LOCK PIN

3 VALVE LOCKING−

HOLE

2 NUT

1

Page: 405Figure 203 Engine Anti−Ice Duct and Valve

TABLE OF CONTENTS

A320 71−80V2500JARB1

Page iFRA US/T-5 Köhler Mar 2006

ATA 71 POWER PLANT 1 . . . . . . . . . . . . . . . . . . . . . . .

ATA 71-00 INTRODUCTION 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE MARK NUMBERS 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE MARK NUMBERS 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTRODUCTION 4 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE DESCRIPTION 6 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 71-00 ENGINE HAZARD AREAS 8 . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 73 ENGINE FUEL AND CONTROL 10 . . . . . . . . . . . . . . . . . . . . . . . . . . . 73−20 FADEC PRESENTATION 10 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC FUNCTIONS 12 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE CONTROL P/B’S AND SWITCHES 14 . . . . . . . . . . . . . . . . . . . . .

ATA 77 INDICATING 18 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77−00 ENGINE INDICATING PRESENTATION 18 . . . . . . . . . . . . . . . . . . .

ATA 72 ENGINE 24 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72-00 ENGINE PRESENTATION 24 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FRONT BEARING COMPARTMENT 26 . . . . . . . . . . . . . . . . . . . . . . . . . . . . NO 4 BEARING COMPARTMENT 28 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . REAR BEARING COMPARTMENT 30 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE MODULES 32 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MODULE 32 INTERMEDIATE CASE 34 . . . . . . . . . . . . . . . . . . . . . . . . . . . . MODULE 31 ( FAN MODULE ) 36 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INLET CONE REMOVAL 38 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FAN BLADE REMOVAL / INSTALLATION 40 . . . . . . . . . . . . . . . . . . . . . . .

ATA 72-31-11 FAN BLADE REPAIR 42 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FAN BLADE INSPECTION / REPAIR 42 . . . . . . . . . . . . . . . . . . . . . . . . . . . PROCEDURE 42 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PROCEDURE 44 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PROCEDURE 46 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MODULE 40 HP COMPRESSOR 48 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMBUSTION SECTION 50 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HP TURBINE 52 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 TH. STAGE MAKE UP AIR VALVE 54 . . . . . . . . . . . . . . . . . . . . . . . . . . . COMMON NOZZLE ASSEMBLY (CNA) 56 . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 72−60 ACCESSORY DRIVE GEARBOX 58 . . . . . . . . . . . . . . . . . . . . . . . ANGLE AND MAIN GEARBOX 58 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DRIVE SEAL 60 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE FLANGES 64 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 72-00 BORESCOPING 66 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GENERAL 66 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . BORESCOOPING GENERAL 68 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . BORESCOOPE INSPECTION OF THE HP COMP. 68 . . . . . . . . . . . . . . . BORESCOPE INSPECTION OF THE HP COMP. CONT. 70 . . . . . . . . . .

ATA 71 POWER PLANT 72 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71-20 ENGINE MOUNTS 72 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71-20 ENGINE MOUNTS 74 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FORWARD ENGINE MOUNT 74 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AFT ENGINE MOUNT 74 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 71-10 NACELLE ACCESS DOORS & OPENINGS 76 . . . . . . . . . . . . . . NACELLE GENERAL 76 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ACCESS DOORS & OPENINGS 76 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FAN COWLS OPENING / CLOSING 78 . . . . . . . . . . . . . . . . . . . . . . . . . . . . FAN COWL LATCH ADJUSTMENT 80 . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 78-32 THRUST REVERSER COWL DOORS 82 . . . . . . . . . . . . . . . . . . . T/R COWLING ( ”C-DUCT” ) OPENING / CLOSING 82 . . . . . . . . . . . . . . THRUST REVERSER HALF LATCHES 84 . . . . . . . . . . . . . . . . . . . . . . . . . LATCH ACCESS PANEL & TAKE UP DEVICE 86 . . . . . . . . . . . . . . . . . . . FRONT LATCH AND OPEN INDICATOR 88 . . . . . . . . . . . . . . . . . . . . . . . . C - DUCT OPENING / CLOSING SYSTEM 90 . . . . . . . . . . . . . . . . . . . . . . C - DUCT HOLD OPEN STRUTS 92 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 79 OIL 94 79−00 GENERAL 94 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79−00 GENERAL 96 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79-30 OIL INDICATING SYSTEM 98 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ECAM OIL INDICATIONS 98 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OIL QUANTITY INDICATING 100 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OIL TEMPERATURE INDICATION 100 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OIL PRESSURE INDICATION 100 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

TABLE OF CONTENTS

A320 71−80V2500JARB1

Page iiFRA US/T-5 Köhler Mar 2006

LOW OIL PRESSURE SWITCH 100 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SCAV. FILT. DIFF. PRESSURE WARNING 100 . . . . . . . . . . . . . . . . . . . . . . . NO.4 BEARING WARNING 100 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OIL TANK 102 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE OIL SERVICING 102 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79-00 OIL SYSYSTEM COMPONENTS 104 . . . . . . . . . . . . . . . . . . . . . . . . . OIL QUANTITY TRANSMITTER 104 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OIL PRESSURE PUMP 106 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AIR COOLED OIL COOLER (ACOC) 108 . . . . . . . . . . . . . . . . . . . . . . . . . . . ACOC OIL TEMPERATURE THERMOCOUPLE 108 . . . . . . . . . . . . . . . . . . FUEL COOLED OIL COOLER (FCOC) 110 . . . . . . . . . . . . . . . . . . . . . . . . . . SCAVENGE SYSTEM 112 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SCAVENGE PUMPS 112 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SCAVENGE OIL COMPONENTS 114 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DE-OILER 116 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NO4 BEARING SCAVENGE VALVE 118 . . . . . . . . . . . . . . . . . . . . . . . . . . . . NO 4 BEARING PRESSURE TRANSDUCER 118 . . . . . . . . . . . . . . . . . . . . NO4 BEAR. SCAV. VALVE DESCRIPTION 120 . . . . . . . . . . . . . . . . . . . . . . NO.4 BEARING SCAVENGE VALVE INDICATING 120 . . . . . . . . . . . . . . . . ENGINE OIL PRESSURE 122 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OIL SYSTEM PRESSURE SENSING 124 . . . . . . . . . . . . . . . . . . . . . . . . . . . LOW OIL PRESSURE SWITCH 124 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MAGNETIC CHIP DETECTORS (M.C.D.) 126 . . . . . . . . . . . . . . . . . . . . . . . MASTER MAGNETIC CHIP DETECTOR 128 . . . . . . . . . . . . . . . . . . . . . . . . IDG OIL SERVICING 130 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 73 ENGINE FUEL AND CONTROL 132 . . . . . . . . . . . . . . . . . . . . . . . . . . . 73−00 FUEL SYSTEM PRESENTATION 132 . . . . . . . . . . . . . . . . . . . . . . . . 73−00 FUEL SYSTEM PRESENTATION 134 . . . . . . . . . . . . . . . . . . . . . . . . DESCRIPTION AND OPERATION 134 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 73−30 INDICATING 136 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GENERAL 136 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL PUMP 138 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL METERING UNIT 138 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 73−10 FUEL DISTRIBUTION COMPONENTS 140 . . . . . . . . . . . . . . . . .

FUEL FILTER 140 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL FILTER DIFF. PRESS. SWITCH 140 . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL TEMPERATURE THERMOCOUPLE 140 . . . . . . . . . . . . . . . . . . . . . . FUEL DIVERTER & RETURN VALVE 140 . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL DISTRIBUTION VALVE 142 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL MANIFOLD AND TUBES 144 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL NOZZLE 144 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IDG FUEL COOLED OIL COOLER 146 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IDG OIL COOLER TEMP. THERMOCOUPLE 146 . . . . . . . . . . . . . . . . . . . . FUEL METERING UNIT 148 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HP & LP FUEL SOV CONTROL 150 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LOW PRESSURE FUEL SHUT OFF VALVE 152 . . . . . . . . . . . . . . . . . . . . .

ATA 73-20 HEAT MANAGEMENT SYSTEM 154 . . . . . . . . . . . . . . . . . . . . . . . . PRESENTATION 154 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL TEMP. THERMOCOUPLE 154 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IDG OIL COOLER TEMP. THERMOCOUPLE 154 . . . . . . . . . . . . . . . . . . . . ACOC OIL TEMP. THERMOCOUPLE 154 . . . . . . . . . . . . . . . . . . . . . . . . . . . ACOC MODULATING AIR VALVE 154 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL DIVERTER & RETURN VALVE 156 . . . . . . . . . . . . . . . . . . . . . . . . . . RETURN TO TANK MODES 156 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HMS MODE 1 ( NORMAL MODE ) 156 . . . . . . . . . . . . . . . . . . . . . . . . . . . . HMS MODE 4 156 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NO RETURN TO TANK MODES 3 AND 5 158 . . . . . . . . . . . . . . . . . . . . . . . HMS MODE 3 158 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HMS MODE 5 158 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AIR MODULATING VALVE 160 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 71-70 POWER PLANT DRAINS 162 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GENERAL 162 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PYLON DRAINS 164 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DRAIN SYSTEM DESCRIPTION 166 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 76 ENGINE CONTROLS 168 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . THROTTLE CONTROL SYSTEM 168 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . THRUST LEVERS 168 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . BUMP RATING PUSH BUTTON 170 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

TABLE OF CONTENTS

A320 71−80V2500JARB1

Page iiiFRA US/T-5 Köhler Mar 2006

ARTIFICIAL FEEL UNIT ( MECANICAL BOX ) 172 . . . . . . . . . . . . . . . . . . . THROTTLE CONTROL UNIT 174 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RIGGING 176 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AIDS ALPHA CALL UP OF TRA 176 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AIDS ALPHA CALL UP OF TRA 178 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 77 INDICATING 180 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77−00 ENGINE INDICATING PRESENTATION 180 . . . . . . . . . . . . . . . . . . .

ATA 77−10 POWER INDICATING 182 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EPR INDICATION 182 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EPR SYSTEM COMPONENTS 184 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . P2 / T2 SENSOR 184 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . P4.9 SENSORS 184 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . P2 / T2 HEATER 186 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC P2/T2 HEATER TEST 188 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 77−20 TEMPERATURE 190 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EGT INDICATION 190 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EGT PROBES 192 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 77−10 POWER 194 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . N1 AND N2 INDICATION 194 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 31 INDICATING 196 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MAX POINTER RESET ( N1, N2 & EGT ) 196 . . . . . . . . . . . . . . . . . . . . . . .

ATA 77-10 POWER 198 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . N1 INDICATION 198 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INTERCHANGE OF N1 SPEED SENSORS 198 . . . . . . . . . . . . . . . . . . . . . . DEDICATED ALTERNATOR (PMA) 200 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VIBRATION INDICATION 202 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE VIBRATION MONITORING UNIT (EVMU) 204 . . . . . . . . . . . . . . . COMPONENTS 206 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EVMU OPERATION (CFDS) 208 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CFDS SYSTEM REPORT / TEST 210 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CFDS SYSTEM REPORT /TEST 212 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CFDS SYSTEM REPORT /TEST 214 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CFDS ACCELEROMETER RECONFIG. 216 . . . . . . . . . . . . . . . . . . . . . . . .

AIRCRAFT INTEGRATED DATA SYSTEM 218 . . . . . . . . . . . . . . . . . . . . . .

ATA 73 ENGINE FUEL AND CONTROL 220 . . . . . . . . . . . . . . . . . . . . . . . . . . . 73−20 FADEC 220 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC LRU‘S 222 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DATA ENTRY PLUG MODIFICATION 222 . . . . . . . . . . . . . . . . . . . . . . . . . . . ELECTRONIC ENGINE CONTROL (EEC) 224 . . . . . . . . . . . . . . . . . . . . . . . FADEC POWER SUPPLY 226 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 73-22 FADEC SENSORS 230 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC LRU‘S SENSORS 230 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC LRU‘S SENSORS 232 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . P3/T3 SENSOR 232 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . P12.5 SENSOR 234 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . P2.5 / T2.5 SENSORS 234 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC DESCRIPTION 236 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC DESCRIPTION 238 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC SYSTEM MAINTENANCE 238 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FAILURES AND REDUNDANCY 240 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FAILURES AND REDUNDANCY 242 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE LIMITS PROTECTION 242 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . POWER MANAGEMENT 244 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AUTOTHRUST ACTIVATION / DEACTIVATION 244 . . . . . . . . . . . . . . . . . . EPR SETTING REQUIREMENTS 250 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RATED N1 SETTING REQUIREMENTS 250 . . . . . . . . . . . . . . . . . . . . . . . . . UNRATED N1 SETTING REQUIREMENTS 250 . . . . . . . . . . . . . . . . . . . . . . IDLE CONTROL 252 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . N1 SPEED TABLE 254 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC FAULT STRATEGY 256 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPONENT FAIL SAFE STATES 258 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LOSS OF INPUTS FROM AIRCRAFT 259 . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 73-20 FADEC TEST 260 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC PREVIOUS LEGS REPORT 260 . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC TROUBLESHOOTING REPORT 262 . . . . . . . . . . . . . . . . . . . . . . . . FADEC FAILURE TYPES DEFINITION 262 . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC SYSTEM TEST 266 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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FADEC GROUND SCANNING 268 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC CLASS 3 FAULT REPORT 268 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC CLASS 3 FAULT REPORT 270 . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 73-25 ENGINE INTERFACE UNIT 272 . . . . . . . . . . . . . . . . . . . . . . . . . . . . EIU PRESENTATION 272 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EIU INPUT DESCRIPTION 272 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EIU INTERFACES 274 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EIU INTERFACES CONT. 275 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CFDS SYSTEM REPORT/TEST EIU 276 . . . . . . . . . . . . . . . . . . . . . . . . . . . LAST LEG REPORT 278 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LRU INDENTIFICATION 278 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GROUND SCANNING 280 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EIU CFDS DISCRETE OUTPUTS SIMULATION 282 . . . . . . . . . . . . . . . . . EIU CFDS DISCRETE OUTPUTS SIMULATION 284 . . . . . . . . . . . . . . . . . EIU DISCRETE OUTPUTS 286 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EIU DISCRETE OUTPUTS 288 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 75 ENGINE AIR 290 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75−00 SYSTEM PRESENTATION 290 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TURBINE COOLING CONTROL 292 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPERATING SCHEDULE 292 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HPT / LPT ACTIVE CLEARANCE CONT. SYS. 294 . . . . . . . . . . . . . . . . . . HPT / LPT COOLING MANIFOLDS 294 . . . . . . . . . . . . . . . . . . . . . . . . . . . . HPT / LPT COOLING MANIFOLDS 296 . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPRESSOR CONTROL 298 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 75-31 LP COMP.AIR FLOW SYS. 300 . . . . . . . . . . . . . . . . . . . . . . . . . . . . BOOSTER BLEED SYSTEM 300 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . BSBV ACTUATING MECHANISM 302 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 75-32 HP COMP. AIR FLOW SYS. 304 . . . . . . . . . . . . . . . . . . . . . . . . . . . . VSV SYSTEM COMPONENTS 304 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VSV RIGGING 306 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HANDLING BLEED VALVES 308 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HANDLING BLEED VALVES 310 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HANDLING BLEED VALVES FUNKTION 312 . . . . . . . . . . . . . . . . . . . . . . . . HANDLING BLEED VALVE MALFUNCTIONS 314 . . . . . . . . . . . . . . . . . . . .

BLEED VALVE LOCATIONS 316 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HANDLING BLEED VALVE MALFUNCTIONS 318 . . . . . . . . . . . . . . . . . . . . NACELLE VENTILATION 322 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 75-41 NACELLE TEMPERATURE 324 . . . . . . . . . . . . . . . . . . . . . . . . . . . . NACELLE TEMPERATURE GENERAL 324 . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 74 IGNITION 326 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74-00 IGNITION SYSTEM PRESENTATION 326 . . . . . . . . . . . . . . . . . . . . . IGNITION SYSTEM COMPONENTS 326 . . . . . . . . . . . . . . . . . . . . . . . . . . . . IGNITION / STARTING− OPERATION 328 . . . . . . . . . . . . . . . . . . . . . . . . . . IGNITION SYSTEM CIRCUIT BREAKERS 328 . . . . . . . . . . . . . . . . . . . . . . IGNITION SYSTEM TEST 330 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IGNITOR TEST 332 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IGNITION SYSTEM TEST 334 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IGNITOR TEST 336 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IGNITION TEST WITHOUT CFDS 338 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 80 STARTING 340 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . GENERAL 340 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . STARTING COMPONENTS 342 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . STARTER AIR CONTROL VALVE 344 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . STARTER AIR CONTROL VALVE 346 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . START AIR CONTROL VALVE TEST 348 . . . . . . . . . . . . . . . . . . . . . . . . . . . START AIR CONTROL VALVE TEST ( FAULT DETECTED ) 350 . . . . . . . CRANKING−DESCRIPTION 352 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . WET CRANKING 354 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AUTOMATIC START 356 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EEC AUTO START ABBORT 356 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MANUAL START 358 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 78 EXHAUST 360 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . REVERSER SYSTEM 360 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . THRUST REVERSER SYSTEM DESCRIPTION 362 . . . . . . . . . . . . . . . . . . THRUST REVERSER INDEPENDENT LOCKING SYSTEM 364 . . . . . . . THRUST REVERSER SYSTEM 366 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . THRUST REVERSER HYDRAULIC SUPPLY 368 . . . . . . . . . . . . . . . . . . . . THRUST REVERSER MANUAL DEPLOYMENT 368 . . . . . . . . . . . . . . . . .

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THRUST REVERSER INDEPENDENT LOCKING SYSTEM 370 . . . . . . . REVERSER HYDRAULIC CONTROL UNIT 372 . . . . . . . . . . . . . . . . . . . . . . HCU IN FORWARD THRUST POSITION 374 . . . . . . . . . . . . . . . . . . . . . . . . HCU DEPLOY SEQUENCE DESCRIPTION 376 . . . . . . . . . . . . . . . . . . . . . HCU STOW SEQUENCE DESCRIPTION 378 . . . . . . . . . . . . . . . . . . . . . . . HYDRAULIC ACTUATION SYS. COMP. 380 . . . . . . . . . . . . . . . . . . . . . . . . . FLEXSHAFT INSTALLATION 380 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HYDRAULIC ACTUATORS DESCRIPTION 382 . . . . . . . . . . . . . . . . . . . . . . UPPER NONLOCKING ACTUATOR 382 . . . . . . . . . . . . . . . . . . . . . . . . . . . . LOWER LOCKING ACTUATORS 384 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . THRUST REVERSER MANUAL DEPLOY / STOW 386 . . . . . . . . . . . . . . . THRUST REVERSER DEACTIVATION 388 . . . . . . . . . . . . . . . . . . . . . . . . . FADEC CFDS REVERSER TEST 390 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FADEC T/R TEST ( FAULT DETECTED ) 392 . . . . . . . . . . . . . . . . . . . . . . . . FADEC T/R TEST ( NOT O.K. ) 394 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 71-00 ENGINE CHANGE 396 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE REMOVAL / INSTALLATION 396 . . . . . . . . . . . . . . . . . . . . . . . . . . .

398 . . . . . . . . . . . POWER PLANT PRESERVATION 398 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ATA 30 ICE AND RAIN PROTECTION 400 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30−20 ENG. AIR INTAKE ICE PROTETION 400 . . . . . . . . . . . . . . . . . . . . . SYSTEM CONTROL 400 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SYSTEM CONTROL SCHEMATIC 402 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ENGINE ANTI ICE DUCT AND VALVE 404 . . . . . . . . . . . . . . . . . . . . . . . . . . ANTI−ICE VALVE DEACTIVATION 404 . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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Figure 1 V2500 Propulsion Unit 5 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 2 Propulsion Unit Outline 7 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 3 Engine Hazard Areas 9 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 4 FADEC Presentation IAE V2500 11 . . . . . . . . . . . . . . . . . . . . Figure 5 FADEC Presentation IAE V2500 13 . . . . . . . . . . . . . . . . . . . . Figure 6 Engine Control P / B‘s and Switches 15 . . . . . . . . . . . . . . . . . Figure 7 Engine Circuit Breakers 16 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 8 Engine Circuit Breakers 17 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 9 Engine ECAM Indications 19 . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 10 Stage Numbering 21 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 11 Engine Stations 23 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 12 Engine Bearings & Compartments 25 . . . . . . . . . . . . . . . . . . Figure 13 Front Bearing Compartment 27 . . . . . . . . . . . . . . . . . . . . . . . Figure 14 No.4 Bearing Compartment 29 . . . . . . . . . . . . . . . . . . . . . . . . Figure 15 Rear Bearing Compartment 31 . . . . . . . . . . . . . . . . . . . . . . . Figure 16 Engine Modules 33 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 17 Fan Case Section 35 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 18 LP Compressor ( Fan ) 37 . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 19 Inlet Cone Removal 39 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 20 Fan Blade Removal / Installation 41 . . . . . . . . . . . . . . . . . . . Figure 21 Fan Blade Repair Limits 43 . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 22 Fan Blade Repair Limits 45 . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 23 Fan Blade Repair Limits 47 . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 24 HP Compressor 49 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 25 Combustion Section 51 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 26 No.4 Bearing Scavenge Valve 53 . . . . . . . . . . . . . . . . . . . . . Figure 27 Stage10 to HPT Air Control Valve 55 . . . . . . . . . . . . . . . . . . Figure 28 Common Nozzle Assemply 57 . . . . . . . . . . . . . . . . . . . . . . . . Figure 29 Angle and Main Gearbox 59 . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 30 Drive Seals 61 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 31 Engine Components Location (L/H side) 62 . . . . . . . . . . . . . Figure 32 Engine Components Location (R/H side) 63 . . . . . . . . . . . . Figure 33 Engine Flanges 65 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 34 Manual Handcranking 67 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 35 HP Compressor Borescope Access 69 . . . . . . . . . . . . . . . . .

Figure 36 HP Compressor Borescope Access 71 . . . . . . . . . . . . . . . . . Figure 37 Mounts and Loads 73 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 38 Engine Mounts 75 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 39 Nacelle Access Doors 77 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 40 Fan Cowls Opening / Closing 79 . . . . . . . . . . . . . . . . . . . . . . Figure 41 Fan Cowl Latch Adjustment 81 . . . . . . . . . . . . . . . . . . . . . . . Figure 42 C-Duct Opening/Closing 83 . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 43 Thrust Reverser Half Latches 85 . . . . . . . . . . . . . . . . . . . . . . Figure 44 Latch Panel & Take Up Device 87 . . . . . . . . . . . . . . . . . . . . . Figure 45 Front Latch with Open Indicator 89 . . . . . . . . . . . . . . . . . . . . Figure 46 ”C” Duct Opening/Closing 91 . . . . . . . . . . . . . . . . . . . . . . . . . Figure 47 „C“ Duct Hold Open Struts 93 . . . . . . . . . . . . . . . . . . . . . . . . Figure 48 Oil System Basic Schematic 95 . . . . . . . . . . . . . . . . . . . . . . . Figure 49 Oil System Schematic 97 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 50 ECAM Oil Indication 99 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 51 Basic Schematic 101 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 52 Oil Tank 103 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 53 Oil Tank 105 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 54 Pressure Pump & Filter 107 . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 55 ACOC Air Flow 109 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 56 Fuel Cooled Oil Cooler 111 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 57 Scavenge Pump Assembly 113 . . . . . . . . . . . . . . . . . . . . . . . . Figure 58 Scavenge Filter,Delta P.Sw and Oil Temp. Sensor 115 . . . . Figure 59 De-Oiler 117 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 60 No.4 Bearing Scavenge Valve 119 . . . . . . . . . . . . . . . . . . . . . Figure 61 No.4 Bearing Scavenge Valve 121 . . . . . . . . . . . . . . . . . . . . . Figure 62 Oil Pressure Chart 123 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 63 LOP Switch and Oil Press. Transmitter 125 . . . . . . . . . . . . . . Figure 64 Chip Detectors 127 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 65 Master Magnetic Chip Detector 129 . . . . . . . . . . . . . . . . . . . . Figure 66 IDG Oil Servicing 131 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 67 Fuel System Schematic 133 . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 68 Fuel System Schematic 135 . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 69 Fuel System Indication 137 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 70 Fuel Pump and Fuel Metering Unit 139 . . . . . . . . . . . . . . . . .

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Figure 71 Fuel Filter Diff. Press. Switch/FCOC Fuel Temp. Thermocouple . 141

Figure 72 Fuel Distribution Valve 143 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 73 Fuel Distribution Tubes 145 . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 74 IDG Fuel Cooled Oil Cooler 147 . . . . . . . . . . . . . . . . . . . . . . . . Figure 75 Fuel Metering Unit Schematic 149 . . . . . . . . . . . . . . . . . . . . . . Figure 76 HP and LP Fuel Shutoff Valve ( SOV ) 151 . . . . . . . . . . . . . . Figure 77 LP Fuel Shut−Off Valve 153 . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 78 HMS Main System Components 155 . . . . . . . . . . . . . . . . . . . Figure 79 Return to Tank Modes 1 and 4 157 . . . . . . . . . . . . . . . . . . . Figure 80 NO Return to Tank Modes 3 and 5 159 . . . . . . . . . . . . . . . . Figure 81 Air Modulating Valve 161 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 82 Drain System 163 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 83 Pylon Drains 165 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 84 Drain System Leakage Test & Limits 167 . . . . . . . . . . . . . . . . Figure 85 Engine Thrust Lever Control 169 . . . . . . . . . . . . . . . . . . . . . . . Figure 86 Bump Push Bottons 171 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 87 Mechanical Boxes 173 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 88 Thrust Control Units 175 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 89 Thrust Control System Rigging 177 . . . . . . . . . . . . . . . . . . . . . Figure 90 Alpha Call−up TRA 179 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 91 Engine ECAM Indications 181 . . . . . . . . . . . . . . . . . . . . . . . . . Figure 92 EPR Indication − Upper ECAM Display Unit 183 . . . . . . . . Figure 93 P2 / T2 and P4.9 Sensor 185 . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 94 P2/T2 Heater Schematic 187 . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 95 P2/T2 Heater Test 189 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 96 EGT Indication 191 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 97 EGT System 193 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 98 N1 and N2 Speed Indication 195 . . . . . . . . . . . . . . . . . . . . . . . Figure 99 Max Pointer Reset 197 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 100 Fan Speed & Trim Balance Sensor,N1 Terminal Block 199 Figure 101 Engine Dedicated Alternator 201 . . . . . . . . . . . . . . . . . . . . . . Figure 102 Vibration Indication 203 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 103 EVMU Schematic 205 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 104 Vibration Sensors 207 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 105 EVMU CFDS Pages 209 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 106 CFDS System Report / Test EVMU 211 . . . . . . . . . . . . . . . . Figure 107 Unbalance Data 213 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 108 Unbalance Data 215 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 109 Accelerometer Reconfiguration 217 . . . . . . . . . . . . . . . . . . . Figure 110 AIDS 219 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 111 FADEC Architecture 221 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 112 EEC/ Data Entry Plug 223 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 113 Electronic Engine Control ( EEC ) 225 . . . . . . . . . . . . . . . . . Figure 114 FADEC Power Supply 227 . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 115 Engine Circuit Breakers 228 . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 116 Engine Circuit Breakers 229 . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 117 FADEC Sensors 231 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 118 P3/T3 Sensor 233 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 119 P2.5 / T2.5 Sensors 235 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 120 FADEC Architecture 237 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 121 FADEC Architecture 239 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 122 FADEC Processing and Fault Logic 241 . . . . . . . . . . . . . . . Figure 123 FADEC Processing and Fault Logic 243 . . . . . . . . . . . . . . . Figure 124 Thrust Control Architecture 245 . . . . . . . . . . . . . . . . . . . . . . . Figure 125 Auto Thrust Defenition 247 . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 126 Thrust Lever Positions 249 . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 127 Power Setting Requirements Schematic 251 . . . . . . . . . . . . Figure 128 Idle Control Requirements 253 . . . . . . . . . . . . . . . . . . . . . . . . Figure 129 Ground Idle Speed Diagram N2 255 . . . . . . . . . . . . . . . . . . . Figure 130 FADEC Single Input Signal Failure 257 . . . . . . . . . . . . . . . . Figure 131 Previous Legs Report 261 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 132 Trouble Shooting Report 263 . . . . . . . . . . . . . . . . . . . . . . . . . Figure 133 Flight Data / Ground Data 264 . . . . . . . . . . . . . . . . . . . . . . . . Figure 134 Flight Data / Ground Data 265 . . . . . . . . . . . . . . . . . . . . . . . . Figure 135 FADEC Self Test 267 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 136 Ground Scanning 269 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 137 FADEC Class 3 Fault Report 271 . . . . . . . . . . . . . . . . . . . . . Figure 138 EIU Schematic 273 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 139 EIU Menu 277 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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Figure 140 Last Leg Rep./ LRU Indentification 279 . . . . . . . . . . . . . . . . Figure 141 Ground Scanning 281 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 142 Discrete Outputs Simulation 283 . . . . . . . . . . . . . . . . . . . . . . Figure 143 Discrete Outputs Simulation 285 . . . . . . . . . . . . . . . . . . . . . . Figure 144 EIU Discrete Outputs 287 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 145 EIU Discrete Outputs 289 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 146 Air Systems Schematic 291 . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 147 Turbine Cooling Control Schematic 293 . . . . . . . . . . . . . . . Figure 148 LPT / HPT Active Clearance Control Valve 295 . . . . . . . . . Figure 149 HPT / LPT Cooling Manifolds 297 . . . . . . . . . . . . . . . . . . . . Figure 150 Compressor Control Schematic 299 . . . . . . . . . . . . . . . . . . . Figure 151 Booster Stage Bleed Valve System 301 . . . . . . . . . . . . . . . . Figure 152 BSBV and Actuating Mechanism 303 . . . . . . . . . . . . . . . . . . Figure 153 VSV System Components 305 . . . . . . . . . . . . . . . . . . . . . . . . Figure 154 VSV Actuator Rig 307 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 155 HP Compressor Bleed Valves 309 . . . . . . . . . . . . . . . . . . . . . Figure 156 HP Compressor Bleed Valves 311 . . . . . . . . . . . . . . . . . . . . . Figure 157 HBV OPEN/CLOSED Schematic 313 . . . . . . . . . . . . . . . . . . Figure 158 Bleed Control Valve Solenoids 315 . . . . . . . . . . . . . . . . . . . . Figure 159 Bleed Valve Locations 317 . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 160 HDLG Bleed Valves Malfunction Tables 319 . . . . . . . . . . . . Figure 161 Bleed Valve Functional Test 320 . . . . . . . . . . . . . . . . . . . . . . Figure 162 Bleed Valve Functional Test(cont) 321 . . . . . . . . . . . . . . . . . Figure 163 Nacelle Ventilation 323 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 164 Nacelle Temperature System 325 . . . . . . . . . . . . . . . . . . . . . Figure 165 Ignition System Components 327 . . . . . . . . . . . . . . . . . . . . . Figure 166 Ignition and Starting System Eng. 1 329 . . . . . . . . . . . . . . . Figure 167 FADEC Ignition Test 331 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 168 FADEC Ignition Test Cont. 333 . . . . . . . . . . . . . . . . . . . . . . . Figure 169 FADEC Ignition Test 335 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 170 FADEC Ignition Test Cont. 337 . . . . . . . . . . . . . . . . . . . . . . . Figure 171 Ignition Test without CFDS 339 . . . . . . . . . . . . . . . . . . . . . . . Figure 172 Starting System Schematic 341 . . . . . . . . . . . . . . . . . . . . . . . Figure 173 Starter Motor 343 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 174 Starter Air Control Valve 345 . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 175 Starter Air Control Valve 347 . . . . . . . . . . . . . . . . . . . . . . . . . Figure 176 Starter Valve Test via CFDS 349 . . . . . . . . . . . . . . . . . . . . . . Figure 177 Starter Valve Test via CFDS 351 . . . . . . . . . . . . . . . . . . . . . . Figure 178 Dry Cranking Procedure 353 . . . . . . . . . . . . . . . . . . . . . . . . . Figure 179 Wet Cranking Procedure 355 . . . . . . . . . . . . . . . . . . . . . . . . . Figure 180 Automatic Start Procedure 357 . . . . . . . . . . . . . . . . . . . . . . . Figure 181 Manual Start Procedure 359 . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 182 Thrust Reverser stowed / deployed 361 . . . . . . . . . . . . . . . . Figure 183 Reverser System Schematic 363 . . . . . . . . . . . . . . . . . . . . . . Figure 184 Reverser System Schematic 365 . . . . . . . . . . . . . . . . . . . . . . Figure 185 Reverser Installation 367 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 186 Reverser Hydraulic Supply 369 . . . . . . . . . . . . . . . . . . . . . . . Figure 187 T/R Independent Locking System (**On A/C 116−199) 371 Figure 188 Hydraulic Control Unit ( HCU ) 373 . . . . . . . . . . . . . . . . . . . . Figure 189 HCU Schematic 375 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 190 HCU Deploy Sequence 377 . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 191 HCU Stow Sequence 379 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 192 Flexible Drive Shafts 381 . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 193 Upper Nonlocking Actuator 383 . . . . . . . . . . . . . . . . . . . . . . . Figure 194 Lower Locking Actuator 385 . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 195 Reverser Manual Operation 387 . . . . . . . . . . . . . . . . . . . . . . Figure 196 T/R Deactivation 389 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 197 FADEC T/R Test (NO FAULT) 391 . . . . . . . . . . . . . . . . . . . . Figure 198 FADEC T/R Test (FAULT DETECTED) 393 . . . . . . . . . . . . . Figure 199 FADEC T/R Test (NOT O.K.) 395 . . . . . . . . . . . . . . . . . . . . . Figure 200 Engine Removal / Installation 397 . . . . . . . . . . . . . . . . . . . . . Figure 201 Engine Nacelle A/I Architecture 401 . . . . . . . . . . . . . . . . . . . Figure 202 Control Schematic 403 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 203 Engine Anti−Ice Duct and Valve 405 . . . . . . . . . . . . . . . . . . .

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