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Preface

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Notice to Holders

The information in this document is the property ofInternational Aero Engines AG and may not be copied, orcommunicated to a third party, or used, for any purpose otherthan that for which it is supplied without the express writtenconsent of International Aero Engines AG.

Whilst this information is given in good faith, based upon thelatest information available to International Aero Engines AG,no warranty or representation is given concerning suchinformation, which must not be taken as establishing anycontractual or other commitment binding International AeroEngines AG or any of its subsidiary or associated companies.

This training manual is not an official publication and must notbe used for operating or maintaining the equipment hereindescribed. The official publications and manuals must be usedfor those purposes: they may also be used for up-dating the

contents of the course notes.

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V2500 ABBREIVATIONS

ACAC Air Cooled Air CoolerACC Active Clearance Control

ACOC Air Cooled Oil Cooler

AIDRS Air Data Inertial Reference System

Alt Altitude

APU Auxiliary Power Unit

AMM Aircraft Maintenance Manual

BDC Bottom Dead Centre

BMC Bleed Monitoring Computer

BSBV Booster Stage Bleed Valve

CFDIU Centralised Fault Display Interface Unit

CFDS Centralised Fault Display System

CL Climb

CNA Common Nozzle Assembly

CRT Cathode Ray Tube

DCU Directional Control Unit

DCV Directional Control Valve

DEP Data Entry Plug

DMC Display Management Computer

ECAM Electronic Centralised Aircraft Monitoring

ECS Environmental Control System

EEC Electronic Engine Control

EGT Exhaust Gas TemperatureEHSV Electro-hydraulic Servo Valve

EIU Engine Interface Unit

EIS Entered Into Service

EVMS Engine Vibration Monitoring System

EVMU Engine Vibration Monitoring Unit

EPR Engine Pressure Ratio

ETOPS Extended Twin Engine Operations

FADEC Full Authority Digital Electronic Control

FAV Fan Air Valve

FCOC Fuel Cooled Oil Cooler

FCU Flight Control Unit

FDRV Fuel Diverter and Return to Tank Valve

FSN Fuel Spray Nozzle

FMGC Flight Management and Guidance Computer

FMV Fuel Metering Valve

FMU Fuel Metering Unit

FOB Fuel On Board

FWC Flight Warning Computer

HCU Hydraulic Control Unit

HIV Hydraulic Isolation Valve

HEIU High Energy Ignition Unit (igniter box)

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HP High Pressure

HPC High Pressure Compressor

HPT High Pressure Turbine

HPRV High Pressure Regulating Valve

HT High Tension (ignition lead)

IDG Integrated Drive Generator

IAE International Aero Engines

IDG Integrated Drive Generator

IFSD In-flight Shut Down

IGV Inlet Guide Vane

lbs. Pounds

LE Leading Edge

LGCIU Landing Gear and Interface Unit

LGCU Landing Gear Control Unit

LH Left Hand

LP Low Pressure

LPC Low Pressure Compressor

LPCBV Low Pressure Compressor Bleed ValveLPSOV Low Pressure Shut off Valve

LPT Low Pressure Turbine

LRU Line Replaceable Unit

LT Low Tension

LVDT Linear Voltage Differential Transformer

MCD Magnetic Chip Detector

MCDU Multipurpose Control and Display Unit

MCLB Max Climb

MCT Max Continuous

Mn Mach Number

MS Micro Switch

NAC Nacelle

NGV Nozzle Guide Vane

NRV Non-Return Valve

N1 Low Pressure system speed

N2 High Pressure system speed

OAT Outside Air Temperature

OGV Outlet Guide Vane

OP Open

OPV Over Pressure Valve

OS Overspeed

Pamb Pressure Ambient

Pb Burner PressurePRSOV Pressure Regulating Shut Off Valve

PRV Pressure Regulating Valve

PSI Pounds Per Square Inch

PSID Pounds Per Square Inch Differential

PMA Permanent Magnet Alternator

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P2 Pressure of the fan inlet

P2.5 Pressure of the LP compressor outlet

P3 Pressure of the HP compressor outlet

P4.9 Pressure of the LP turbine outlet

QAD Quick Attach/Detach

SAT Static Air Temperature

SEC Spoiler Elevator Computer

STS Status

TAI Thermal Anti Ice

TAT Throttle Angle Transducer

TAP Transient Acoustic Propagation

TCT Temperature Controlling Thermostat

TDC Top Dead Centre

TE Trailing Edge

TEC Turbine Exhaust Case

TFU Transient Fuel Unit

TRA Throttle Resolver Angle

TLA Throttle Lever AngleTLT Temperature Limiting Thermostat

TM Torque Motor

TO Take-off

TOBI Tangential out Board Injector

TX Transmitter

UDP Uni-directionally Profiled

VIGV Variable Inlet Guide Vane

VSV Variable Stator Vane

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V2500 GENERAL FAMILARISATION COURSE NOTES CONTENTS

PREFACE

SECTION 1 ENGINE INTRODUCTION

SECTION 2 PROPULSION SYSTEM, FIRE PROTECTION AND VENTILATION

SECTION 3 ENGINE MECHANICAL ARRANGEMENT

SECTION 4 ELECTRONIC ENGINE CONTROL

SECTION 5 POWER MANAGEMENT

SECTION 6 FUEL SYSTEM

SECTION 7 OIL SYSTEM

SECTION 8 HEAT MANAGEMENT SYSTEM

SECTION 9 COMPRESSOR AIRFLOW CONTROL SYSTEM

SECTION 10 SECONDARY AIR SYSTEMS

SECTION 11 ENGINE ANTI-ICE SYSTEM

SECTION 12 ENGINE INDICATATIONS

SECTION 13 STARTING AND IGNITION SYSTEM

SECTION 14 THRUST REVERSE

SECTION 15 TROUBLESHOOTING

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INTRODUCTION

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IAE International Aero Engines AG 2000

IAE V2500 General Familiarisation Introduction

IAE V2500 Line and Base Maintenance for Engineers

This is not an Official Publication and must not be used for

operating and maintaining the equipment herein described.The Official Publications and Manuals must be used forthese purposes.

These course notes are arranged in the sequence ofinstruction adopted at the Rolls Royce Customer TrainingCentre.

Considerable effort is made to ensure these notes areclear, concise, correct and up to date. Thus reflectingcurrent production standard engines at the date of the lastrevision.

The masters are updated continuously, but copies areprinted in economic batches. We welcome suggestions forimprovement, and although we hope there are no errors orserious omissions please inform us if you discover any.

Telephone:

Outside the United Kingdom (+44) 1332 - 244350

Within the United Kingdom 01332 244350

Your instructor for this course is:

----------------------------------------------------------------------------

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IAE International Aero Engines AG (IAE)

On March 11, 1983, five of the worlds leading aerospace

manufacturers signed a 30 year collaboration agreementto produce an engine for the single isle aircraft market withthe best proven technology that each could provide. Thefive organisations were:

Rolls Royce plc - United Kingdom.

Pratt and Whitney - USA.

Japanese Aero Engines Corporation.

MTU-Germany.

Fiat Aviazione -Italy.

In December of the same year the collaboration wasincorporated in Zurich, Switzerland, as IAE InternationalAero Engines AG, a management company established to

direct the entire program for the shareholders.The headquarters for IAE were set up in East Hartford,Connecticut, USA and the V2500 turbofan engine to powerthe 120-180 seat aircraft was launched on January 1

st

1984.

Each of the shareholder companies was given theresponsibility for developing and delivering one of the five

engine modules. They are:

Rolls Royce plc - High Pressure Compressor.

Pratt and Whitney Combustion Chamber and High

Pressure Turbine.

Japanese Aero engine Corporation (JAEC) - Fan andLow Pressure Compressor.

Motoren Turbinen Union (MTU) - Low PressureTurbine.

Fiat Aviazione - External Gearbox.

Note: Rolls Royce have developed and introduced thewide chord fan to the V2500 engine family.

The senior partners Rolls Royce and Pratt and Whitneyassemble the engines at their respective plants in DerbyEngland and Middletown Connecticut USA. IAE is

responsible for the co-ordination of the manufacture andassembly of the engines. IAE is also responsible for thesales, marketing and in service support of the V2500.

Note:Fiat Aviazione have since withdrawn as a risk-sharing partner, but still remains as a Primary Supplier.Rolls Royce now has responsibility for all external gearbox

related activity.

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IAE V2500 Engine/Airframe ApplicationsThe V2500 engine has been designated the V because

International Aero Engines (IAE) was originally a five-nation consortium. The V is the Roman numeral for five.

The 2500 numbering indicated the first engine type to bereleased into production. This engine was rated at25000lbs of thrust.

For ease of identification of the present and all futurevariants of the V2500, IAE has introduced an engine

designation system.

All engines possess the V2500 numbering as a genericname.

The first three characters of the full designation areV25. This will identify all the engines in the family.

The next two figures indicate the engines rated sea

level takeoff thrust.

The following letter shows the aircrafts manufacturer.

The last figure represents the mechanical standard ofthe engine.

This system will provide a clear designation of a particularengine as well as a simple way of grouping by name

engines with similar characteristics.

The designation V2500-D collectively describes all

applications for the Boeing McDonnell Douglas MD-90aircraft.

The V2500-A collectively describes all the applicationsfor the Airbus Industries aircraft.

This is irrespective of engine thrust rating.

The number given after the alpha indicates the mechanical

standard of the engine. For example;

V2527-A5.

The only engine exempt from these idents is the currentservice engine, which is already certified to the designatedV2500-A1. There is only one standard of this engine ratingand is utilised on the Airbus A320 aircraft.

Note:

The D5 variant is now no longer in production, howeverthe engine is still extensively overhauled and re-furbished.

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Introduction to the Propulsion System

The V2500 family of engines share a common design

feature for the propulsion system.

The complete propulsion system comprises the engineand the nacelle. The major components of the nacelle areas follows:

The intake cowl.

The fan cowl doors.

Hinged C- ducts with integral thrust reverser units.

Common nozzle assembly.

Intake Cowl

The pitot style inlet cowl permits the efficient intake of airto the engine whilst minimising nacelle drag.

The intake cowl contains the minimum of accessories. Thetwo main accessories that are within the intake cowl are:

P2/T2 probe.

Thermal anti icing ducting and manifold.

Fan Cowl Doors

Access to the units mounted on the fan case and external

gearbox can be gained easily by opening the hinged fancowling doors.

The fan cowl doors are hinged to the aircraft pylon in fourpositions.

There are four quick release adjustable latches thatsecure the fan cowl doors in the closed position.

Each fan cowl doors has two integral support struts that

are secured to the fan case to hold the fan cowl doors inthe open position.

C - Duct Thrust Reverser units

The C-ducts is hinged to the aircraft pylon at fourpositions per C-duct and is secured in the closed positionby six latches located in five positions.

The C-ducts is held in the open position by two integralsupport struts.

Opening of the C-ducts allows access to the core engine.

Common Nozzle Assembly (CNA)

The CNA exhausts both the fan stream and core enginegas flow through a common propulsive nozzle.

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Engine

The V2500 is a twin spool, axial flow, and high bypass

ratio turbofan type engine.The engine incorporates several advanced technologyfeatures, which include:

Full Authority Digital Electronic Control (FADEC).

Wide chord fan blades.

Single crystal HP turbine blades.

'Powdered Metal' HP turbine discs.

A two-piece, annular combustion system, which utilisessegmental liners.

Engine Mechanical Arrangement

The low-pressure (LP) system comprises a single stagefan and multiple stage booster. The booster, which is

linked to the fan, has:

A5 standard four stages.

A1 standard three stages.

The boosters are axial flow type compressors.

A five-stage LP turbine drives the fan and booster.

The booster stage has an additional feature. This is anannular bleed valve, which has been incorporated toimprove starting and handling.

Three bearing assemblies support the LP system. Theyare:

A single ball type bearing (thrust).

Two roller type bearings (support).

The HP system comprises of a ten-stage axial flowcompressor, which is driven by a two-stage HP turbine.The HP compressor has variable inlet guide vanes (VIGV)and variable stator vanes (VSV).

The A5 standard has one stage of VIGV and threestages of VSVs.

The A1 standard has one stage of VIGV and fourstages of VSV's.

The HP system utilises four bleed air valves. These valvesare designed to bleed air from the compressors so as toimprove both starting and engine operation and handling

characteristics.

Two bearing assemblies support the HP system. They are:

A single ball type bearing (thrust).

A single roller type bearing (support).

The combustion system is of an annular design,constructed with an inner and outer section.

There are twenty fuel spray nozzles supplying fuel to thecombustor. The fuel is metered according to the setting ofthe thrust lever or the thrust management computer via theFADEC system.

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The FADEC system uses pressures and temperatures ofthe engine to control the various systems for satisfactoryengine operation. The sampling areas are identified as

stations and are common to all variants of the V2500engine.

The following are the measurement stations for the V2500engine:

Station 1 - Intake/Engine inlet interface.

Station 2 - Fan inlet.

Station 2.5 LPC Outlet Guide Vane (OGV) exit.

Station 12.5 - Fan exit/ C-Duct by-pass air.

Station 3 - HP Compressor exit.

Station 4.9 - LP Turbine exit.

Engine stage numbering

The V2500 engine has compressor blade numbering asfollows:

Stage 1 - Fan.

Stage 1.5 - LPC booster

Stage 2 - LPC booster.

Stage 2.3 - LPC booster (A5 Only).Stage 2.5 - LPC booster.

Stages (3-12) - HPC Stages.

Note the HPC is a ten-stage compressor.

The V2500 engine has turbine blade stage numbering asfollows:

Stages (1-2) - HP Turbine Stages.

Stages (3-7) - LP Turbine Stages.

V2500-A1 V2527-A5

EIS May 89 Dec 93

Take-off thrust (lb) 25,000 26,500

Flat rate temp (C) 30 45

Fan diameter (ins) 63 63.5

Airflow (lb/s) 792 811

Bypass ratio 5.4 4.8

Climb-pressure ratio 35.8 32.8

Cruise sf (lbf/lb/hr) 0.543 0.543

Power plant wt. (lb) 7400 7500


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SECTION 2

PROPULSION SYSTEM, FIRE PROTECTION

&

VENTILATION


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IAE V2500 General Familiarisation Propulsion System, Fire Protection and Ventilation

Propulsion System Introduction

Purpose

The propulsion system encloses the Powerplant. Theyprovide the ducting for the fan bypass air and provide foran aerodynamic exterior.

Description

The propulsion system comprises of the engine and thefollowing nacelle units:

Intake cowl assembly.

The L and R hand hinged fan cowl doors.

The thrust reverser C-ducts.

The common nozzle assembly (CNA).

Engine mounts for the front and rear of the engine.

Fire protection and ventilation system.

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Airframe Interfaces

Purpose

The airframe interfaces provide a link between the engineand aircraft systems.

Description

The following units form the interface between the aircraftand engine:

The front and rear engine mounts.

The bleed air off-takes.

The starter motor air supply.

Integrated Drive Generator (IDG) electrical power.

Fuel supplies.

Hydraulic fluid supplies.

FADEC system interfaces.

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Propulsion System Access Panels

Purpose

The propulsion system access panels provide the engineerwith quick access to the components that require regularor scheduled inspection.

The access panels allow the removal and installation ofLine Replaceable Units (LRUs) during maintenanceactivities.

DescriptionThe access panels provided on the propulsion system areas follows:

Engine Left Hand Side

Fan cowl door

Oil tank service door.

Master magnetic chip detector panel.

Zone 1 Ventilation Outlet Grille for the Fan Case.

Thrust reverser C-duct

Maintenance access panels for the thrust reverserhydraulic actuators.

Translating cowl lockout pins.

Engine Right Hand Side

Intake cowl

Interphone jack.

Anti icing outlet grille.

P2/T2 probe access panel.

Fan cowl doors

Air-cooled oil cooler outlet.

Starter motor air valve access panel.

Zone 1 Ventilation Outlet Grille for the Fan Case.

Breathers overboard discharge.

Thrust reverser C duct

Maintenance access panels for the thrust reverserhydraulic actuators.

Translating cowl lockout pins.

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Propulsion System Core Engine Access

Purpose

The propulsion system can be opened to allow access forengineers both to the fan case and core engine.

Description

Fan cowl doors

The fan cowl doors are hinged from the aircraft strut at thetop and are secured by four latches at the bottom.

When in the open position they are supported by twosupport struts per Fan Cowl.

Thrust reverser C ducts

The Thrust Reverser C-ducts are hinged from the aircraftstrut at the top by four hinged type brackets and aresecured by six latches at the bottom.

When in the open position they are supported by twosupport struts per C-duct.

Propulsion System Materials and Weights

Intake cowl

The intake cowl is made up of the following materials:

Intake D section is aluminium.

Intake cowl is carbon fibre.

Intake cowl weight is 238 lbs. (107.98 Kg).

Fan cowl doors

The fan cowl doors are made up of the following materials:

Carbon fibre and aluminium.

LH fan cowl door weight is 79 lbs. (35.84 Kg).

RH fan cowl door weight is 86 lbs. (39.01 Kg).

Thrust Reverser C-ducts

The thrust reverser C ducts are made up of the followingmaterials:

C-duct structure and translating cowls are carbon fibreand aluminium.

The thrust reverser C-duct weight is 578 lbs. (262.25Kg).

Common nozzle assembly (CNA)The CNA is made up of the following material:

Titanium.

CAN weight is 213 lbs.

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Intake Cowl

PurposeTo supply all the air required by the engine, with minimumpressure losses and with an even pressure face to the fan.

Nacelle drag is also minimised due to the aerodynamicallystreamlined design.

Location

The inlet cowl is bolted to the front of the LPC case (Fan).

Description

The intake cowl is constructed from hollow inner and outerskins. These are supported by front (titanium) and rear(Graphite/Epoxy composite) bulkheads.

Inner and outer skins are manufactured from composites.

The leading edge is a 'one piece' pressing in Aluminium.The cowl weight is approximately 238 lbs.

The intake cowl has the following features:

Integral thermal anti-icing system.

P2T2 Probe.

Ventilation Intake. Interphone socket.

Engine attachment ring with alignment pins to ensurecorrect location of the cowl on to the fan case.

Door locators that automatically align the fan cowl doorsto ensure good sealing.

Strut brackets to provide location for the left and righthand fan cowl door support struts (front struts only).

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Fan Cowl Doors (FCD)

PurposeThe fan cowl doors provide for an aerosmooth exterior whileenclosing the fan case mounted accessories.

Location

They are located about the fan casing.

Hinged at the top to the aircraft strut and secured by four

latches at the bottom.

Description

The doors extend rearwards from the inlet cowl to overlapleading edge of the 'C' ducts.

The A320 aircraft have a strake on the inboard cowl of eachengine, the right hand cowl on both engine 1 and left-hand

cowl on engine 2.The A319 aircraft have strakes on both the left-hand and righthand cowls on both engines 1 and 2.

Fan cowls are interchangeable between the A319 and A320except for the strake configuration. Make sure the correctconfiguration is installed.

The fan cowl doors are constructed from graphite skinsenclosing an aluminium honeycomb inner.

Aluminium is also used to reinforcement each corner tominimises handling/impact damage and wear.

The fan cowl doors abut along the bottom centre line and aresecured to each other by 4 quick release and adjustablelatches.



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Thrust Reverser C Ducts

PurposeThe thrust reverser C ducts provide for:

An aerosmooth exterior to minimise drag.

The fan bypass ducting.

Reverse thrust for aircraft deceleration.

Location

The thrust reverser C ducts are hinged from the aircraftstrut at the top and are secured at the bottom by six toggletype clamps.

Description

The thrust reverser C ducts extend rearwards from the fancowls to the common nozzle assembly (CNA).

The thrust reverser C ducts;

Form the cowling around the core engine (inner barrel) toassist in stiffening the core engine (load-share).

Form the fan air duct between the fan case exit and theentrance to the CNA.

House the thrust reverser operating mechanism and

cascades.Form the outer cowling between the fan cowl doors andCNA.

The thrust reverser C ducts are mostly constructed fromcomposites but some sections are metallic mainlyaluminium for example the inner barrel, blocker doors andlinks.

The thrust reverser C-ducts can be opened for access tothe core engine. This allows maintenance to be carried outon the core engine while the engine is installed to theaircraft.

The thrust reverser C-ducts are heavy therefor hydraulicactuation is required to open them. Normal engine oil isused in a hand-operated pump.

The thrust reverser C-ducts are held in the open positionby two support struts.

The forward strut is a fixed length.

The rear strut is a telescopic support.



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Combined Nozzle Assembly (CNA)

Purpose

The CNA allows the mixing of the hot and cold stream gasflows to produce the resultant thrust.

Location

The CNA is bolted to the rear flange of the turbine exhaustcasing. There is no fixing to the bottom of the pylon.

Description

The CNA:

Forms the exhaust unit.

Mixes the hot and cold gas streams and ejects thecombined flow to atmosphere through a singlepropelling nozzle.

Completes the engine nacelle.



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Engine Mounts

Purpose

The engine mounts suspend the engine from the aircraftstrut.

The engine mounts transmit loads generated by theengine during aircraft operation.

Location

The front engine mount is located at the rear of the

intermediate case at the core engine.The rear engine mount is located on the LPT casing atTDC.

Description

Forward engine mount

The forward engine mount is designed to transmit thefollowing loads:

Thrust loads.

Side loads.

Vertical loads.

The front mount is secured to the intermediate case inthree positions:

A monoball type universal joint. This gives the mainsupport at the front engine mount position.

Two thrust links that are attached to:

The cross beam of the engine mount.

Support brackets either side of the monoball location.

Rear engine mount

The rear engine mount is designed to transmit the

following loads: Torsional loads.

Side loads.

Vertical loads.

The rear engine mount has a diagonal main link that givesresistance to torsional movement of the casing as a result

of the hot gas passing through the turbines.There is further support from two side links. These limit theengine side to side movement and give vertical support.



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FIRE PROTECTION AND VENTILATION



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Fire Protection and Ventilation

Purpose

The purpose of fire protection is to give an indication to theflight deck of a possible fire condition about the engine.

The purpose of the ventilation system is to provide a flowof cooling air about the engine to reduce the risk of a firecondition annunciation to the flight deck.

Location

The locations of the fire detection fire wires are about thefan casing and core engine.

The location of the ventilation air is about the entire of thefan case and core engine.

Description

The engine is ventilated to provide a cooling airflow for

maintaining the engine components within an acceptableoperating temperature.

Also to provide a flow of air that assists in the removal ofpotential combustible liquids that may be in the area.

Ventilation is provided for:

The fan case area (Zone 1).

The core engine area (Zone 2).Zones 1 and 2 are ventilated to:

Prevent accessory and component over heating.

Prevent the accumulation of flammable vapours.

Zone 1 ventilation

Ram air enters the zone through an inlet located on the

upper LH side of the air intake cowl.The air circulates through the fan compartment and exitsat the exhaust located on the bottom rear centre line of thefan cowl doors.

Zone 2 ventilation

Metered holes within the inner barrel of the C duct allowpressurized fan air to enter the zone 2 area.

Air exhausting from the active clearance control (ACC)system around the turbine area also provides ventilationair for Zone 2.

The air circulates through the core compartment and exitsthrough the lower bifurcation of the C ducts via the thrustrecovery duct.

Ventilation during ground running

During ground running local pockets of natural convectionexist providing some ventilation of the fancase zone 1.

Zone 2 ventilation is provided by the fan duct pressure asabove during ground running and in flight.




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Fire Detection System

Purpose

The fire detection system monitors the air temperature inZone 1 and Zone 2.

When the air temperature increases to a pre determinedlevel the system provides flight deck warning.

Location

The fire detection system is located:

Routed around the high-speed external gearbox.

At BDC of the core engine nearest to the combustordiffuser case.

Description

The V2500 utilises a Systron Donner fire detection system.It has a gas filled core and relies upon heat exposure to

increase the internal gas pressure. Thus triggeringsensors.

When the air temperature about the fan case and/or coreengine increases to a pre-determined level the system isdesigned to detect this and display a warning messageand indications to the flight deck.

The system provides flight deck warning by:

Master warning light.

Audible warning tone.

Specific ECAM fire indications.

Engine fire push button illuminates.

Zone 1 and Zone 2 fire detectors function independently ofeach other.

Each zone has two detector units which are mounted as apair, each unit gives an output signal when a fire oroverheat condition occurs.

The two detector units are attached to support tubes byclips.

Nacelle air temperature (NAC)

Zone 2 has the nacelle air temperature sensor.

Indication is to the flight deck when a temperatureexceedance has occurred.




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Fi D t ti S t d D t t U it Fi i d t t

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Fire Detection System and Detector Units

The fire detection system employs detector units calledfirewires.

The firewires are mounted in pairs. This is necessary dueto the class1 level 3 message that they generate when afire or overheat condition exists.

The fire detection system comprises of the following units:

The firewires send a signal to the Fire Detection Unit(FDU).

The FDU sends a signal to the Flight WarningComputer (FWC).

The FWC generates the flight deck indications for a firecondition.

There is one FDU per engine. The FDU has two channels;each channel is looking at a separate fire detector loop ofzones 1 and 2.

Under normal conditions both firewires require to beindicating to the FDU to give a real indication to the flightdeck.

If there is a single loop failure of more than 16 secondsthen the remaining firewire will continue to operate. TheFDU will recognise the faulty fire loop.

The faulty loop will be indicated to ECAM as the followingmessage:

ENG 1 (2) FIRE LOOP A (B) FAULT

If there is a double loop failure then the FDU will recognisethis as a possible burn through and the fire message willbe generated to the flight deck.

Firewire detectors

Each of the fire wire detector units comprises of thefollowing:

A hollow sensor tube.

A responder assembly.

Sensor tube

The sensor tube is closed and sealed at one end and theother open end is connected to the responder.

The tube is filled with helium gas and carries a central coreof ceramic material impregnated with hydrogen.

An increase in the air temperature around the sensor tubecauses the helium to expand and increase until thepressure causes the alarm switch to close. The FDUrecognises this as an abnormal situation, hence fireindication will be illuminated.

If a burn through occurs, the pressure within the sensingtube is lost and as a result of this the integrity switchopens to give an indication to the FDU of a loop failure.

Responder

The responder has two pressure switches, one normallyopen and the other normally closed.

The normally open switch is the alarm indication.

The normally closed switch is the fault indication.




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Fire Detection System Fire Bottles

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Fire Detection System Fire Bottles

Purpose

The fire bottles provide a means of extinguishing apotentially hazardous fire about the engine when a fireannunciation to the flight deck has occurred.

Location

The engine fire bottles are located in the aircraft strut.Access for maintenance is via a panel that can be foundon the left hand side.

Description

The fire bottles have the following features:

Agent type is bromotrifluoromethane.

Charged to a nominal pressure of 600 psi at 21 deg. C.

Pressure switch.

Discharge head.

Discharge squibs.

The pressure switch is set to indicate bottle empty whenthe pressure falls below 225 psi. The indication in the flightdeck is:

AGENT 1 (2) SQUIB DISC

This is an illuminating annunciator light on the overheadpanel.

The discharge head has a leak proof diaphragm that isdesigned to rupture when:

The squib is activated from the flight deck.

Excessive pressure in the fire bottle. 1600 to 1800 psiat 95 deg. C

The squib is an Electro Pyrotechnic Cartridge containing

explosive powder. Two filaments ignite the powder whenthey are supplied with 28v dc.

There is facility to carry out a fire system test that will giveall the expected indications if all is functioning correctly.

The fire test switch is located on the fire push button panelon the overhead panel.




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Fire Detection System Indications and Controls

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Fire Detection SystemIndications and Controls

Purpose

The purpose of the fire detection system indications is toalert the flight crew to a possible fire condition.

The controls allow the flight crew to react and deal with theimpending fire indication in the flight deck.

Location

The fire control panel is located on the overhead panel forfire bottle operation and fire system test.




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Nacelle Air Temperature (NAC)

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p ( )

Purpose

The nacelle air temperature gives an advisory indication tothe lower ECAM CRT if a temperature exceedance hasbeen experienced.

Location

The NAC sensor is located by the bifurcation panel atbottom dead centre between the two-thrust reverser Cduct halves.

The NAC is in zone 2.

Description

Under normal conditions the NAC indication is notdisplayed on the lower ECAM CRT.

When a temperature exceedance of 320 deg.c hasoccurred the indication will appear to the lower ECAMCRT.

This indication is displayed if;

The system is not in engine starting mode and one of thetwo temperatures reaches the advisory threshold.




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SECTION 3

MECHANICAL ARRANGEMENT


IAE V2500 General Familiarisation Mechanical Arrangement

Mechanical Arrangement General

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The engine is an axial flow, high by-pass ratio, twin spoolturbo fan.

The general arrangement is shown below.

L.P. System

Four stage L.P. compressor - comprising:

1 Fan stage

L.P. Compressor consisting of 4 stages driven by:

Five stage L.P. Turbine

H.P. System

Ten-stage axial flow compressor driven by a 2 stageH.P. Turbine.

Variable angle inlet guide vanes.

Variable stator vanes (3 stages A5).

Handling bleed valves at stage 7 and 10.

Customer service bleeds at stage 7 and 10

Combustion System

Annular, two piece, with 20 fuel spray nozzles.

Gearbox Radial drive via a tower shaft from H.P. Compressor

shaft to fan case mounted Angle and Main gearboxes.

Gearbox provides mountings and drive for all enginedriven accessories and the pneumatic starter motor.

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Engine Main Bearings

Th i b i t d th b i b i

No 4 Bearing

R di l t f t bi d f H P h ft

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The main bearing arrangement and the bearing numberingsystem is shown below.

The 5 bearings are located in 3 bearing compartments:

The Front Bearing Compartment, located at the centreof the Intermediate Case, houses No's 1,2 & 3 bearings.

The Centre Bearing Compartment located in thediffuser/combustor case houses No 4 Bearing.

The Rear Bearing Compartment located in the Turbine

Exhaust Case houses No 5 Bearing.

No 1 Bearing

Shaft axial location bearing.

Takes the thrust loads of the L.P. shaft.

Single track ball bearing.

No 2 Bearing Radial support for the front of the L.P.turbine shaft.Single track roller bearing utilising "squeeze film" oildamping.

No 3 Bearing

H.P. shaft axial location bearing.

Radial support for the front of the H.P.shaft. Takes the thrust loads of the H.P. shaft.

Single track ball bearing.

Mounted in a hydraulic damper, which is centred by aseries of rod springs (squirrel cage).

Radial support for turbine end of H.P. shaft.

Single track roller bearing.No 5 Bearing

Radial support for the turbine end of the L.P. shaft.

Single track roller bearing.

Squeeze film oil damping.

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Engine Internal Cooling and Sealing Airflows Cooling, sealing and scavenge air for the No.4 Bearing

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Engine Internal Cooling and Sealing AirflowsPurpose

To provide sealing air for the bearing chambers so as toprevent oil loss.

To provide cooling air for the engines internal componentskeeping them within designed operating temperatures.

Location

The air used for internal cooling and sealing is taken from thecompressor stages of:

LPC stage 2.5

HPC stage 8.

HPC stage 10.

HPC stage 12. The fan bypass provides external cooling air.

Description

Fan air is used to provide:

Air for the Active Clearance Control (ACC) system.This is used to control the tip clearances of the turbineblades.

Air through the Air Cooled Air Cooler (ACAC). This isused for the precooling of the buffer air.

Buffer air is used to provide:

Cooling, sealing and scavenge air for the No.4 BearingChamber.

LPC stage 2.5 air is used for

Sealing of the front and rear of the Front BearingChamber

HPC stage 7 air is used for airflow control for compressorstability and aircraft services bleed supply.

HPC stage 8 is used for:

Sealing the hydraulic seal of the Front BearingChamber and the sealing of the No. 5 BearingChamber.

HPC stage 10 air is used for:

Airflow control and aircraft services supply.

Make up air supply for the HPT stage 2 disc andblades.

Cooling air for the HPT stage 2 NGVs.

HPC stage 12 air is used for:

Combustion chamber cooling.

HPT stage 1 blades and NGVs cooling.

The supply to the ACAC for buffer air cooling andsealing of the no. 4 bearing chamber.

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Modular Construction

Modular construction has the following advantages:

Note:

The module numbers refer to the ATA chapter reference

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Modular construction has the following advantages:

lower overall maintenance costs maximum life achieved from each module

reduced turn-around time for engine repair

reduced spare engine holdings

ease of transportation and storage

rapid module change with minimum ground running easy hot section inspection

vertical/horizontal build strip

split engine transportation

compressors/turbines independently balanced

Module DesignationModule No Module

31 Fan

32 Intermediate

40 HP System

41 - HP Compressor

45 - HP Turbine

50 LP Turbine

60 External gearbox

The module numbers refer to the ATA chapter referencefor that module.

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Module 31

Description

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p

Module 31 (Fan Module) is the complete Fan assemblyand comprises:

22 Hollow fan blades

22 Annulus Fillers

Fan Disc

Front and Rear Blade Retaining Rings

The blades are retained in the disc radially by the dovetailroot.

The front and rear blade retaining rings provides axialretention. Blade removal/replacement is easily achieved byremoving the front blade retaining ring and sliding theblade along the dovetail slot in the disc.

22 annulus fillers form the fan inner annulus.The nose cone and fairing smooth the airflow into the fan.

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Module 32 - Intermediate Case

The Intermediate Module comprises of:

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Fan Case Fan Duct

Fan Outlet Guide Vanes (OGV)

LP Compressor ( A5 - 4 stage)

LP Compressor Bleed Valve (LPCBV)

Front engine mount structure Front bearing compartment which houses Nos. 1, 2

and 3 bearings

Drive gear for the power off-take shaft (gearbox drive)

LP stub shaft

Inner support struts

Outer support struts

Vee groove locations for the inner and outer barrels ofthe 'C' ducts




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Module 32 - Intermediate Case

Instrumentation

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The following pressures and temperatures are sensed andtransmitted to the E.E.C.

P12.5

P2.5

T2.5

The rear view of the intermediate case is shown below.




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Module 40 HP Compressor

Description

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The HP compressor assembly (Module 40 is a 10 stageaxial flow compressor. It has a rotor assembly and statorcase. The compressor stages are numbered from thefront, with the first stage is stage being designated asstage 3 of the whole engines compressor system. Airflowthrough the compressor is controlled by variable inletguide vanes (VIGV), variable stator vanes (VSV) andbleed valves.

The rotor assembly has five sub-assemblies

(1) Stages 3 to 8 HP compressor disks

(2) A vortex reducer ring.

(3) Stages 9 to 12 HP compressor disks

(4) The HP compressor shaft.

(5) The HP compressor rotating air seal.

The five sub-assemblies are bolted together to make therotor. The compressor blades in stages 3 to 5 are attachedto the compressor disks in axial dovetail slots and securedby lockplates. The stages 6 to 12 compressor blades areinstalled in slots around the circumference of the disksthrough an axial loading slot. Lock blades, lock nuts and

jack screws hold the blades in position.The HP compressor stator case has two primary sub-assemblies, the HP compressor front and rear cases.




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Module 40 HP Compressor

The HP compressor front case assembly has two split

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cases bolted together along the engine horizontal centreline.

The front case assembly contains the VIGVs, the stages 3to 5 VSVs and the stage 6 stator vanes.

The front outer case provides a mounting for the VIGV andVSV actuator. The front case assembly is bolted to the

intermediate case and to the rear outer case.

The HP compressor rear case assembly has five inner ringcases and an outer case. Flanges on the inner cases formannular manifolds, which provide stages 7 and 10 airofftakes.

The five inner cases are bolted together, with the frontsupport cone bolted at the stage 7 case and the stage 11case bolted to the rear outer case. The five inner casescontain the stages 7 to 11 fixed stator vanes.

The rear outer case is bolted to the diffuser case and tothe rear flange of the HP compressor front case.

Access is provided in the compressor cases for borescopeinspection of the compressor blades and stator vanes




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Combustion Section

The combustion section includes the diffuser section, thecombustion inner and outer liners, and the No 4 bearing

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g

assembly.

Diffuser Casing

The diffuser section is the primary structural part of thecombustion section.

The diffuser section has 20 mounting pads for theinstallation of the fuel spray nozzles. It also has two

mounting pads for the two ignitor plugs.Combustion Liner

The inner and outer liners form the combustion liner.

The outer liner is located by five locating pins, which passthrough the diffuser casing.

The inner combustion liner is attached to the turbine

nozzle guide vane assembly.

The inner and outer liners are manufactured from sheetmetal with 100 separate liner segments attached to theinner surface (50 per inner and outer liner). The segmentscan be replaced independently during engine overhaul.




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HP Turbine

Description

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The primary parts of the HP turbine rotor and statorassembly are:

The HP Turbine Rotor Assemblies (Stage 1 and Stage 2)

The HP Turbine Case and Vane Assembly

The HP turbine rotor assemblies are two stages of turbinehubs with single-crystal, nickel-alloy blades. The two-hub

configuration removes a bolt flange between hubs. Thisdecreases the weight and enables faster engine assembly.

The blades have airfoils with high strength and resistanceto creep. Satisfactory blade tip clearances are supplied byactive clearance control (ACC) to cool the case withcompressor air.




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LP Turbine

Description

The primary parts of the Low Pressure Turbine (LPT) The five LPT disks are made from high heat resistant

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p y p ( )module are:

LPT Five Stage Rotor

LPT Five Stage Stator Vanes

Air Seals

LPT Case

Inner and Outer Duct

LPT Shaft

Turbine Exhaust Case (TEC)

The LP turbine has a five stage rotor which supplies powerto the LP compressor through the LPT shaft. The LPTrotor is installed in the LPT case where it is in alignment

with the LPT stators. The LPT case is made from high-heat resistant nickel alloy and is a one part weldedassembly. To identify the LP turbine module, anidentification plate is attached to the LP turbine case at the136degrees position.

The LPT case has two borescope inspection ports at125.27 and 237.10 degrees. The ports are used to

internally examine the adjacent engine sections:

Trailing Edge (TE), Stage 2, HPT Blades

Leading Edge (LE), Stage 3, LPT Blades

Trailing Edge (TE), Stage 3, LPT Blades

gnickel alloy. The LPT blades are also made from nickelalloy and are attached to the disks by fir-tree roots. Theblades are held in axial position on the disk by the rotatingair seals (knife-edge).




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Module 60 - External Gearbox

Purpose

The gearbox assembly transmits power from the engine to The following accessory units are located on the external

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provide drives for the accessories mounted on the gearboxfront and rear faces.

During engine starting the gearbox also transmits powerfrom the pneumatic starter motor to the core engine.

The gearbox also provides a means of hand cranking theHP rotor for maintenance operations.

Location

The gearbox is mounted by 4 flexible links to the bottom ofthe fan case.

Main gearbox 3 links.

Angle gearbox 1 link.

Description

The external gearbox is a cast aluminium housing that hasthe following features;

Individually replaceable drive units.

Magnetic chip detectors.

Main gearbox 2 magnetic chip detectors.

Angle gearbox 1 magnetic chip detector.

gearbox;

Front Face Mount Pads

De-oiler.

Pneumatic starter.

Dedicated generator.

Hydraulic Pump.

Oil Pressure pump and filter.

Rear Face Mount Pads

Fuel pumps (and fuel metering unit FMU).

Oil scavenge pumps unit.

Integrated drive generator (IDG).

The Oil sealing for the gearbox to accessory drive links isprovided by a combination of carbon and O-ring typeseals.

The carbon seals can be replaced while the engine is onwing.




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Engine View Right Hand Side

The following components are located on the right handside of the engine.

1 St 10 k i l f l t t bi

19. LPT and HPT active clearance control valves (ACC).

20. HPC stage 10 handling bleed valve.

21. Engine rear mount.

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1. Stage 10 make-up air valve for supplementary turbinecooling.

2. IDG harness interface.

3. Harness interface.

4. Start air and anti ice ducting interface.

5. Electrical harness interface.

6. Air starter duct.

7. Engine electronic control.

8. Anti ice duct.

9. Relay box.

10. Anti ice valve.

11. Starter valve.

12. 10thstage handling bleed valve solenoid.

13. No.4 bearing scavenge valve.

14. Air-cooled oil cooler (ACOC).

15. Intergrated drive generator (IDG).

16. Exciter ignition boxes.

17. Fuel distribution valve.

18. HPC stage 7B handling bleed valve.

22. Booster bleed valve slave actuator.

23. Front engine mount.

24. HPC 10thstage cooling air for the HPT 2

ndstage NGVs.

25. Solenoids for the three off HPC 7th stage handling

bleed valves.

26. Solenoid for the HP10 make-up cooling air controlvalve.

27. Solenoid for the HP10 cabin bleed pressureregulating/shut-off valve (PRSOV).




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Engine View Left Hand Side

The following components are located on the left-handside of the engine.

1 Fan cowl door hinged brackets (4 off)

20. HPC 7thstage bleed valve (HPC7 C).

21. HPC 7thstage cabin bleed non-return valve (NRV).

22. VIGV/VSV actuator.

23 F l d f l i i

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1. Fan cowl door hinged brackets (4 off).

2. Thrust reverser hydraulic control valve (HCU).

3. Hydraulic tubes interface.

4. Fuel supply and return to wing tank.

5. C duct front hinge.

6. Thrust reverser hydraulic tubes interface.7. Over pressuerization valve (OPV).

8. 2.5 bleed master actuator.

9. C Duct floating hinges.

10. Fan Air Valve (FAV).

11. C Duct rear hinge.

12. Opening actuator mounting brackets.

13. C Duct compression struts (3off).

14. Cabin bleed air pre cooler duct interface.

15. Cabin bleed air system interface.

16. Pressure regulating valve (PRV).

17. Air-cooled air cooler (ACAC).

18. HPC 10thstage cabin bleed offtake pipe.

19. HPC 10th

stage pressure regulating/shut-off valve(PRSOV).

23. Fuel pumps and fuel metering unit.

24. High speed external gearbox.

25. Hydraulic pump.

26. Engine oil tank.

27. IDG oil cooler.

28. LP fuel filter.

29. Fuel cooled oil cooler (FCOC).

30. Savenge oil filter pressure differential switch.

31. Fuel return to tank valve (part of item 32).

32. Fuel diverter valve (part of item 31).

33. Oil pressure differential transmitter.34.Low oil pressure switch.




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Engine Combined Drains System

Purpose

To provide an early indication of a system or component

failure by evidence of a fluid leak

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failure by evidence of a fluid leak.

Location

The drains systems of tubes are located about the engine.

The drains mast is located at BDC of the fan case. Itprotrudes from the bottom of the fan cowl doors.

Description

This provides a combined overboard drain through adrains mast. The drains are for fuel and oil from the coremodule components, the LP compressor/intermediatecase components and the external gearbox.




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SECTION 4

ELECTRONIC ENGINE CONTROL


IAE V2500 General Familiarisation Electronic Engine Control

Electronic Engine Control Introduction

The V2500 uses a Full Authority Digital Electronic EngineControl (FADEC).

The FADEC comprises the sensors and data input, theelectronic engine control unit (EEC) and the output

Six screened pressure ports provide the requiredpressure inputs to both channels.

Built in handle facilitates removal and handling.

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electronic engine control unit (EEC) and the outputdevices, which include solenoids, fuel servo operatedactuators and pneumatic servo operated devices. TheFADEC also includes electrical harnesses.

Engine Electronic Control

The heart of the FADEC is the Engine Electronic Control

(EEC) unit - shown below. The EEC is a fan case mountedunit, which is shielded and grounded as protection againstEMI - mainly lightning strikes.

Features

Vibration isolation mountings.

Shielded and grounded (lightning strike protection).

Size - 15.9 X 20.1 X 4.4 inches.

Weight - 41 lbs.

Two independent electronic channels.

Two independent power supplies, the EEC utilises67.53 Watts of power from either the three phase AC

from a dedicated engine mounted alternator, or 28Volts DC from an aircraft source.

A two way Pressure Relief Valve maintains the unitsdifferential pressure (< 5 PSID).

Has three control modes in each channel. EnginePressure Ratio (EPR) which is the Primary thrustcontrol Mode. N1 Rated and Un-rated and alsoprovides Auto Starting and Thrust Reverser control. (Tobe covered in detail later).

Schedules engine operation to provide maximumengine performance and fuel savings.

Provides improved engine starting (Auto Start) andtransient characteristics (acceleration/deceleration).

Provides maximum engine protection and is moreflexible to readily adapt to changes in enginerequirements.

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The Engine Electronic Control (EEC) Description

The EEC is a dual channel control unit that utilises a splithousing design.

The assembled unit is sealed with a housing seal and aprotective shield provides channel separation.

Each of the EEC channels can exercise full control of allengine functions. Control alternates between Channel Aand Channel B for consecutive flights, the selection of thecontrolling channel being made automatically by the EEC

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protective shield provides channel separation.

The control assembly is separated into two modules, eachcontaining one control channel.

Each module contains two multi-layer printed circuitboards assemblies, which enable it to functionindependently of the other channel.

A mating connector provides Crosstalk, for partial orcomplete channel switching and fault isolation logic whenthe two modules are joined.

This connector also provides for the exchange of cross-link data, cross wiring and hardwired discretes betweenthe two channels.

The EEC has two identical electronic circuits that areidentified as Channel A and Channel B. Each channel issupplied with identical data from the aircraft and theengine.

This data includes throttle position, aircraft digital data, airpressures, air temperatures, exhaust gas temperaturesand rotor speeds.

The EEC, to set the correct engine rating for the flightconditions uses this data. The EEC also transmits engineperformance data to the aircraft.

This data is used in cockpit display, thrust managementand condition monitoring systems.

itself.

The channel not in control is nominated as the back upchannel

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.

Electronic Engine Control

Harness (electrical) and Pressure Connections

Two identical, but separate electrical harnesses providethe input/output circuits between the EEC and the relevant

Electrical Connections

Front Face

J1 E.B.U. 4000 KSA

J2 Engine D202P

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the input/output circuits between the EEC and the relevantsensor/control actuator, and the aircraft interface.

The harness connectors are 'keyed' to preventmisconnection.

Note: Single pressure signals are directed to pressure

transducers - located within the EEC - the pressuretransducers then supply digital electronic signals tochannels A and B.

The following pressures are sensed: -

Pamb ambient air pressure - fan case sensor

Pb burner pressure (air pressure) P3/T3 probe

P2 fan inlet pressure - P2/T2 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

J3 Engine D203P

J4 Engine D204P

J11 Engine D211P

Rear Face

J5 Engine D205P

J6 Data Entry Plug

J7 E.B.U. 4000 KSB

J8 Engine D208P

J9 Engine D209P

J10 Engine D210P

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Engine Electronic Control (EEC.)

Overview

The EEC provides the following engine control functions:-

Power Setting (E.P.R.).

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Acceleration and deceleration times.

Idle speed governing.

Overspeed limits (N1 and N2).

Fuel flow.

Variable stator vane system (V.S.V.)

Compressor handling bleed valves.

Booster stage bleed valve (B.S.B.V.).

Turbine cooling (10 stage make-up air system).

Active clearance control (A.C.C.).

Thrust reverser.

Automatic engine starting.

Oil and fuel temperature management.

Note:

The fuel cut off (engine shut down) command comes from

the flight crew and is not controlled by the EEC.Fault Monitoring

The EEC has extensive self test and fault isolation logicbuilt in. This logic operates continuously to detect andisolate defects in the EEC.

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Electronic Engine Control (EEC) Data Entry Plug

Purpose

The Data Entry Plug (DEP) provides discrete data inputs

to the EEC. Located on to Junction 6 of the EEC. itprovides unique engine data to Channel A and B. The datat itt d b th DEP i

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transmitted by the DEP is:

EPR Modifier (Used for power setting).

Engine Rating (Selected from multiple rating options).

Engine Serial No.

Location

The data entry plug is located on the channel B sideelectrical connectors of the EEC.

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DATA ENTRY PLUG (DEP)



Electronic Engine Control

Failures and Redundancy

Improved reliability is achieved by utilising dual sensors,

dual control channels, dual selectors and dual feedback. Dual sensors are used to supply all EEC inputs except

( i l t d ithi th EEC

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pressures, (single pressure transducers within the EECprovide signals to each channel - A and B).

The EEC uses identical software in each of the twochannels. Each channel has its own power supply,processor, programme memory and input/output

functions. The mode of operation and the selection ofthe channel in control is decided by the availability ofinput signal and output controls.

Each channel normally uses its own input signals buteach channel can also use input signals from the otherchannel required i.e. if it recognises faulty, or suspect,inputs.

An output fault in one channel will cause switchover tocontrol from the other channel.

In the event of faults in both channels a pre-determinedhierarchy decides which channel is more capable ofcontrol and utilises that channel.

In the event of loss of either channels, or loss of

electrical power, the systems are designed to go to thefail safe positions.




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Operation and Control

EEC Power Supplies

The electrical supplies for the EEC are normally provided

by a dedicated alternator, which is mounted to and drivenby the external gearbox.

Dedicated Alternator

The EEC also utilises aircraft power to operate some

engine systems:- 115 volts AC 400 Hz power is required for the ignition

system and inlet probe anti icing heater

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Dedicated Alternator

The unit is a permanent magnet alternator which has twoindependent sets of stator windings and supplies twoindependent, 3 phase, frequency wild AC outputs to theEEC These unregulated AC supplies are rectified to 28

volts DC within the EEC

The Dedicated Alternator also supplies the N2 (HPCompressor speed) signal for the EEC. This is provided bythe frequency of a single phase winding in the statorhousing. This source is the primary speed signal and isused by both Channels of the EEC and for the Flight Deckinstrument display of engine actual speed. Should this

signal fail, there is a Back-up signal which is derived fromone of the three phase windings of Channel B powergeneration.

There is no speed signal generation provided by the outputof the coil windings of the Dedicated Alternators ChannelA power supply.

system and inlet probe anti-icing heater

28V DC is required for some specific functions, whichinclude the thrust reverser, fuel on/off and ground andtest power for EEC maintenance.

In the event of a dedicated alternator total failure the EECis supplied from the aircraft 28V DC bus bars, 28V DCfrom the same source is also used by the EEC duringengine starts until the dedicated alternator comes 'on line'at approximately 10% N2.

The dedicated alternator comes on line and supplies theEEC power requirement when the N2 reachesapproximately 10%.

Switching between the aircraft 28V supply and dedicatedalternator power supplies is done automatically by theEEC.

The dedicated alternator is cooled by 12.5 cooling air.piped from the fan exit pressure probe, which is mountedin the upper fan case splitter fairing.

.




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SECTION 5

POWER MANAGEMENT


IAE V2500 General Familiarisation Power Management

Power Management

Purpose

The power management system is designed to allow the

control of engine power by either manual or auto throttlecontrol.

Location

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The aircraft throttle is located in the flight deck. This is inreference to the TLA resolvers.

The EEC is engine intermediate case mounted. This is inreference to the TRA signal that is derived from TLA.

Description

The throttle control lever (Thrust Lever) is based on the"fixed throttle" concept, there is no motorised movement ofthe throttle levers.

Each throttle control lever drives dual throttle resolvers,each resolver output is dedicated to one EEC channel.

The throttle lever angle (TLA) is the input to the resolver.

The resolver output, which is fed to the EEC, is known asthe Throttle Resolver Angle (TRA).

The relationship between the throttle lever angle and thethrottle resolver angle is linear therefor;

1 deg TLA = 1.9 deg TRA

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Throttle Control Lever Mechanism

The throttle control mechanism for one engine is shownbelow.

The control system consists of: The throttle control lever.

The mechanical box

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The mechanical box.

The throttle control unit.

The throttle control lever movement is transmitted througha rod to the mechanical box. The mechanical box

incorporates 'soft' detents which provides selected engineratings, it also provides "artificial feel" for the throttlecontrol system.

The output from the mechanical box is transmitted by asecond rod to the throttle control unit. The throttle controlunit incorporates two resolvers and six potentiometers.

Each resolver is dedicated to one EEC. channel, the

output from the potentiometers provides T.L.A. signals tothe aircraft flight management computers.

A rig pin position is provided on the throttle control unit forrigging the resolvers and potentiometers.

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Bump Rating Push Button (A1 Engined Aircraft only)

In some cases (optional) the throttle control levers areprovided with "Bump" rating push buttons, one per engine.This enables the EEC to be re-rated to provide additional

thrust capability for use during specific aircraft operations.

Note:

Bump Ratings can be selected regardless of TLA only in

Flexible Takeoff (A1 & A5Engined Aircraft)

Definition of Flexible Takeoff:

In many instances, the aircraft takes off with a weight

lower than the maximum permissible takeoff weight. Whenthis happens, it can meet the required performance with adecreased thrust that is adapted to the weight: This iscalled Flexible Takeoff and the thrust is called Flexible

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Bump Ratings can be selected, regardless of TLA only inEPR mode when aircraft is on ground.

Bump Ratings can be de-selected at any time by actuatingthe bump rating push button, as long as the aircraft is on

the ground and the Thrust Lever is not in the Max Take-Offdetent.

In flight, the bump ratings are fully removed when theThrust Lever is moved from the Take-Off detent to orbelow the Max Continuous detent.

The Bump Rating is available in flight (EPR or N1 mode)under the following conditions;

Bump Rating is initially selected on ground.

Take-Off, Go Around TOGA Thrust position set.

Aircraft is within the Take-Off envelope.

When Bump Rating is selected a B appears next to theassociated EPR display. Use of Bump must be recorded.

When one Bump button is selected, both engines areBump Rated.

Pressing Bump again deselects Bump Rating.

called Flexible Takeoff and the thrust is called FlexibleTakeoff Thrust. The use of Flexible Takeoff Thrust savesengine life.

The maximum permissible takeoff weight decreases astemperature increases, so it is possible to assume atemperature at which the actual takeoff weight would bethe limiting one. This temperature is called FlexibleTemperature or Assumed Temperature and is enteredinto the FADEC via the MCDU PERF TO page in order toget the adapted thrust.

Note! If the thrust Bump is armed for takeoff and flexiblethrust is used, the pilot must use the Takeoff Performance

determined for the non-increased takeoff thrust (withoutBump).

Thrust must not be reduced by more than 25% of thefull rated thrust.

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Throttle Control Lever Mechanism

The throttle control lever moves over a range of 65degrees, from minus 20 degrees to plus 45 degrees. Anintermediate retractable mechanical stop is provided at 0

degrees.

Forward Thrust Range

The forward thrust range is from 0 degrees to plus 45

Thrust Rating Limit

Thrust rating limit is computed according to the thrust leverposition. If the thrust lever is set in a detent the FADEC willselect the rating limit corresponding to this detent.

If the thrust lever is set between two detents the FADECwill select the rating limit corresponding to the highermode.

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The forward thrust range is from 0 degrees to plus 45degrees.

0 degrees = forward idle power.

45 degrees = rated take off power.

Two detents are provided in this range;

Max climb (MCLB) at 25 degrees.

Max continuous (MCT)/Flexible (de-rated) take offpower (FLTO) at 35 degrees.

Reverse Thrust Range

Lifting the reverse latching lever allows the throttle tooperate in the range 0 degrees to minus 20 degrees. Adetent at minus 6 degrees corresponds to thrust reversedeploy commanded and reverse idle power, minus 20degrees is max reverse power.

Auto Thrust System (ATS)

The Auto Thrust System can only be engaged between 0degrees and plus 35 degrees.

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EEC/Fuel System Interface

Purpose

To allow the throttle signal from the flight deck to bereceived by the EEC. The EEC will convert this signal intoa fuel flow error in order to change the fuel flow for apower level change.

Description

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