V2500 lbm issue 01

Preface

Notice to Holders The information in this document is the property of International Aero Engines AG and may not be copied, or communicated to a third party, or used, for any purpose other than that for which it is supplied without the express written consent of International Aero Engines AG. Whilst this information is given in good faith, based upon the latest information available to International Aero Engines AG, no warranty or representation is given concerning such information, which must not be taken as establishing any contractual or other commitment binding International Aero Engines AG or any of its subsidiary or associated companies. This training manual is not an official publication and must not be used for operating or maintaining the equipment herein described. The official publications and manuals must be used for those purposes: they may also be used for up-dating the contents of the course notes.

V2500 A1/A5 (Airbus A319/320/321) Line & Base Maintenance Course

Time Table

Session 1 Session 2 Session 3 Session 4 Session 5

Day 1 Induction & registration Introduction Propulsion System Engine Mechanical

Arrangement

Day 2 Engine Mechanical Arrangement Fan Maintenance

Day 3 Fan Maintenance FADEC Power Management

Day 4 Fuel System Oil System Heat Management

Day 5 Airflow Control System

V2500 A1/A5 (Airbus A319/320/321) Line & Base Maintenance Course

Time Table

Session 1 Session 2 Session 3 Session 4 Session 5

Day 6

Secondary Air System

Anti-icing System

Engine Systems Indication Starting & Ignition

Day 7 Thrust Reverser On-board Maintenance Systems & Trouble-shooting

Day 8 On-board Maintenance Systems & Trouble-shooting

Examination

V2500 ABBREIVATIONSACAC Air Cooled Air Cooler ACC 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 Temperature EHSV 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)

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 Valve LPSOV 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 Pressure PRSOV Pressure Regulating Shut Off Valve PRV Pressure Regulating Valve PSI Pounds Per Square Inch PSID Pounds Per Square Inch Differential PMA Permanent Magnet Alternator

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 Angle TLT 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

V2500 LINE AND BASE MAINTENANCE COURSE NOTES CONTENTS

PREFACE

SECTION 1 ENGINE INTRODUCTION

SECTION 2 PROPULSION SYSTEM, FIRE PROTECTION AND VENTILATION

SECTION 3 ENGINE MECHANICAL ARRANGEMENT

SECTION 4 FAN BLADE REPLACEMENT & FAN TRIM BALANCE

SECTION 5 ELECTRONIC ENGINE CONTROL

SECTION 6 POWER MANAGEMENT

SECTION 7 FUEL SYSTEM

SECTION 8 OIL SYSTEM

SECTION 9 HEAT MANAGEMENT SYSTEM

SECTION 10 COMPRESSOR AIRFLOW CONTROL SYSTEM

SECTION 11 SECONDARY AIR SYSTEMS

SECTION 12 ENGINE ANTI-ICE SYSTEM

SECTION 13 INSTRUMENTATION

SECTION 14 STARTING AND IGNITION SYSTEM

SECTION 15 THRUST REVERSE

SECTION 16 TROUBLESHOOTING

INTRODUCTION

© IAE International Aero Engines AG 2000 V2500 Line and Base Maintenance 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 for these purposes. These course notes are arranged in the sequence of instruction adopted at the Rolls Royce Customer Training Centre. Considerable effort is made to ensure these notes are clear, concise, correct and up to date. Thus reflecting current production standard engines at the date of the last revision. The masters are updated continuously, but copies are printed in economic batches. We welcome suggestions for improvement, and although we hope there are no errors or serious 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: ----------------------------------------------------------------------------

Revision 2 Page 1-1

© IAE International Aero Engines AG 2000 V2500 Line and Base Maintenance Introduction IAE International Aero Engines AG (IAE) On March 11, 1983, five of the worlds leading aerospace manufacturers signed a 30 year collaboration agreement to produce an engine for the single isle aircraft market with the best proven technology that each could provide. The five 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 was incorporated in Zurich, Switzerland, as IAE International Aero 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 power the 120-180 seat aircraft was launched on January 1st 1984. Each of the shareholder companies was given the responsibility 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 and Low Pressure Compressor.

• Motoren Turbinen Union (MTU) - Low Pressure Turbine.

• Fiat Aviazione - External Gearbox. Note: Rolls Royce have developed and introduced the wide chord fan to the V2500 engine family. The senior partners Rolls Royce and Pratt and Whitney assemble the engines at their respective plants in Derby England and Middletown Connecticut USA. IAE is responsible for the co-ordination of the manufacture and assembly of the engines. IAE is also responsible for the sales, 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.

Revision 2 Page 1-2

© IAE International Aero Engines AG 2000 V2500 Line and Base Maintenance Introduction

Revision 2 Page 1-3

© IAE International Aero Engines AG 2000 V2500 Line and Base Maintenance Introduction IAE V2500 Engine/Airframe Applications The 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 be released into production. This engine was rated at 25000lbs of thrust. For ease of identification of the present and all future variants of the V2500, IAE has introduced an engine designation system.

• All engines possess the V2500 numbering as a generic name.

• The first three characters of the full designation are V25. 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 of the engine.

This system will provide a clear designation of a particular engine 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-90 aircraft.

• The V2500-A collectively describes all the applications for 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 current service engine, which is already certified to the designated V2500-A1. There is only one standard of this engine rating and is utilised on the Airbus A320 aircraft. Note: The D5 variant is now no longer in production, however the engine is still extensively overhauled and re-furbished.

Revision 2 Page 1-4


Revision 2 Page 1-5


THIS PAGE IS LEFT INTENTIONALLY BLANK

Revision 2 Page 1-6

V2

Re

© IAE International Aero Engines AG 2000 500 Line and Base Maintenance Introduction

vision 2 Page 1-7

ENGINE SPECIFICATIONS & APPLICATIONS OF V2500

V2500-A1

V2522-A5

V2524-A5

V2527-A5*

V2530-A5

V2533-A5

V2525-D5

V2528-D5

Application A320 A319 A319 A320 A321 A321 MD-90 MD-90

Engine inService

May 89 Dec 97 Jun 97 Dec 93 Mar 94 Mar 97 Apr 95 Apr 95

Take-off thrust(lb)

25,000 22,000 24,000 26,500 31,400 33,000 25,000 28,000

Flat rate temp. C. 30 55 55 45 30 30 30 30

Fan diameter(ins)

63 63.5 63.5 63.5 63.5 63.5 63.5 63.5

Bypass ratio 5.4 4.9 4.9 4.8 4.6 4.5 4.8 4.7

Cruise sfc(lbf/lb/hr)

0.543 0.543 0.543 0.543 0.543 0.543 0.543 0.543

Powerplant wt(lb)

7,400 7,500 7,500 7,500 7,500 7,500 7,900 7,900

Enhanced version 27E for ‘hot and ‘high’ operators and 27M available forcorporate jet A319 application with increased ‘climb’ rate faciltiy.

© IAE International Aero Engines AG 2000 V2500 Line and Base Maintenance Introduction 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 engine and the nacelle. The major components of the nacelle are as 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 air to the engine whilst minimising nacelle drag. The intake cowl contains the minimum of accessories. The two 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 fan cowling doors. The fan cowl doors are hinged to the aircraft pylon in four positions. There are four quick release – adjustable latches that secure 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 in the open position. C - Duct Thrust Reverser units The ‘C’-ducts is hinged to the aircraft pylon at four positions per ‘C’-duct and is secured in the closed position by six latches located in five positions. The ‘C’-ducts is held in the open position by two integral support 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 engine gas flow through a common propulsive nozzle.

Revision 2 Page 1-8


Revision 2 Page 1-9

© IAE International Aero Engines AG 2000 V2500 Line and Base Maintenance Introduction Engine The V2500 is a twin spool, axial flow, and high bypass ratio turbofan type engine. The engine incorporates several advanced technology features, 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 utilises

segmental liners. Engine Mechanical Arrangement The low-pressure (LP) system comprises a single stage fan 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 an annular bleed valve, which has been incorporated to improve starting and handling.

Three bearing assemblies support the LP system. They are: • A single ball type bearing (thrust). • Two roller type bearings (support). The HP system comprises of a ten-stage axial flow compressor, 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 three stages of VSV’s.

• The A1 standard has one stage of VIGV and four stages of VSV's.

The HP system utilises four bleed air valves. These valves are designed to bleed air from the compressors so as to improve 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 the combustor. The fuel is metered according to the setting of the thrust lever or the thrust management computer via the FADEC system.

Revision 2 Page 1-10



V2500 Line and Base Maintenance Introduction

The FADEC system uses pressures and temperatures of the engine to control the various systems for satisfactory engine operation. The sampling areas are identified as stations and are common to all variants of the V2500 engine. The following are the measurement stations for the V2500 engine:

• 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 as follows: 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 as follows:

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


V2500 Line and Base Maintenance Introduction


SECTION 2

PROPULSION SYSTEM (Chapter 71) FIRE PROTECTION (Chapter 26)

COOLING &VENTILATION (Chapter 75)

© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Propulsion System Introduction Purpose The propulsion system encloses the Powerplant. They provide the ducting for the fan bypass air and provide for an aerodynamic exterior. Description The propulsion system comprises of the engine and the following 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.

Revision 2 Page 2-1

© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System

Revision 2 Page 2-2

© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Airframe Interfaces Purpose The airframe interfaces provide a link between the engine and aircraft systems. Description The following units form the interface between the aircraft and 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.

Revision 2 Page 2-3


Revision 2 Page 2-4

© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Propulsion System Access Panels Purpose The propulsion system access panels provide the engineer with quick access to the components that require regular or scheduled inspection. The access panels allow the removal and installation of Line Replaceable Units (LRU’s) during maintenance activities. Description The access panels provided on the propulsion system are as follows: Engine Left Hand Side Fan cowl door Oil tank servicing panel. Master magnetic chip detector panel. Zone 1 Ventilation Outlet Grille for the Fan Case. Thrust reverser C-duct Maintenance access panels for the thrust reverser hydraulic 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 reverser hydraulic actuators. Translating cowl lockout pins.

Revision 2 Page 2-5


Revision 2 Page 2-6

© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Propulsion System Core Engine Access Purpose The propulsion system can be opened to allow access for engineers both to the fan case and core engine. Description Fan cowl doors The fan cowl doors are hinged from the aircraft strut at the top and are secured by four latches at the bottom. When in the open position they are supported by two support struts per Fan Cowl. Thrust reverser C ducts The Thrust Reverser C-ducts are hinged from the aircraft strut at the top by four hinged type brackets and are secured by six latches at the bottom. When in the open position they are supported by two support 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 93 lbs. (42 Kg).

• RH fan cowl door weight is 105 lbs. (47 Kg). Thrust Reverser C-ducts The thrust reverser C ducts are made up of the following materials;

• C-duct structure and translating cowls are carbon fibre and aluminium.

• The thrust reverser C-duct weight is 561 lbs. (257 Kg). Common nozzle assembly (CNA) The CNA is made up of the following material;

• Titanium.

• The CNA weight is 213 lbs.

Revision 2 Page 2-7


Revision 2 Page 2-8

© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Intake Cowl Purpose To supply all the air required by the engine, with minimum pressure losses and with an even pressure face to the fan. Nacelle drag is also minimised due to the aerodynamically streamlined 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 outer skins. 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 ensure correct location of the cowl on to the fan case.

• Door locators that automatically align the fan cowl doors to ensure good sealing.

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

Revision 2 Page 2-9



© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Fan Cowl Doors (FCD) Purpose The two fan cowl doors provide for an aerodynamically smooth exterior, while enclosing the fan case mounted accessories. Location They are located about the fan casing. Four hinges attach each fan cowl door to the aircraft pylon. Description The doors extend rearwards from the inlet cowl to overlap leading edge of the 'C' ducts. The A320 aircraft have a strake on the inboard cowl of each engine, 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 right hand cowls on both engines 1 and 2. Fan cowls are interchangeable between the A319 and A320 except for the strake configuration. Make sure the correct configuration is installed. The fan cowl doors are constructed from graphite skins enclosing an aluminium honeycomb inner. Aluminium is also used to reinforcement each corner to minimises handling/impact damage and wear. The fan cowl doors abut along the bottom centre line and

are secured to each other by 4 quick release and adjustable latches. Warning The fan cowl hold open struts must be in the extended position and both struts must always be used to hold the doors open. Be careful when opening the doors in winds of more than 26 knots (30 mph). The fan cowl doors must not be opened in winds of more than 52 knots (60 mph). SB V2500-NAC-71-0259 Introduces a device that holds the fan cowl doors in a partial open position when the doors are unsupported by the struts. This device makes clear whether the fan cowl doors are secured closed or are unlatched and unsupported. SB V2500-NAC-71-0227 The latches are coloured orange so as to be easily recognised. They are also designed to hang vertical when they are not latched in the close position.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Thrust Reverser C Ducts Purpose The thrust reverser C ducts provide for; • An aerodynamically smooth exterior to minimise drag. • The fan bypass ducting. • Reverse thrust for aircraft deceleration. Location The thrust reverser C-ducts are hinged from the aircraft strut at the top and are secured at the bottom by six latches. Description The thrust reverser C-ducts extend rearwards from the fan cowls to the common nozzle assembly (CNA). The thrust reverser C ducts; Form the cowling around the core engine (inner barrel) to assist in stiffening the core engine (load-share). Form the fan air duct between the fan case exit and the entrance to the CNA. House the thrust reverser operating mechanism and cascades. Form the outer cowling between the fan cowl doors and CNA. The thrust reverser C-ducts are mostly constructed from composites but some sections are metallic mainly aluminium for example the inner barrel, blocker doors and links.

The thrust reverser C-ducts can be opened for access to the core engine. This allows maintenance to be carried out on the core engine while the engine is installed to the aircraft. The thrust reverser C-ducts are heavy, therefore hydraulic actuation is required to open them. Normal aircraft engine lubrication oil is used in a hand-operated pump. The thrust reverser C-ducts are held in the open position by two support struts.

• The forward strut is a fixed length.

• The rear strut is a telescopic support. Warning Both struts must always be used to support the thrust reverser C-ducts in the open position. The unit weight is approximately 578 lbs each. Serious injury to personnel working under the thrust reverser C-ducts can occur if they are suddenly released. Note: Damage to the hinge access panel (HAP) will occur if the C-ducts are opened with the translating cowl in the deploy position. Damage to the wing leading edge slats will occur if they are in the extended position when opening the C-ducts.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Combined Nozzle Assembly (CNA) Purpose The CNA allows the mixing of the hot and cold stream gas flows to produce the resultant thrust. This mixing of the hot and cold gas streams within the CNA reduces the ‘thermal shear effect’ of the gases exiting the propelling nozzle to atmosphere. Additionally, acoustic properties of the CNA minimise still further the noise levels produced by the gas stream. This system results in the V2500 being one of the quietest engines in its class. An important factor as current and future legislation regarding noise pollution at airports is becoming a major issue. Location The CNA is bolted to the rear flange of the turbine exhaust casing. 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 the combined flow to atmosphere through a single propelling nozzle.

• Completes the engine nacelle.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Engine Mounts Purpose The engine mounts suspend the engine from the aircraft strut. The engine mounts transmit loads generated by the engine 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 at TDC. Description Forward engine mount The forward engine mount is designed to transmit the following loads;

• Thrust loads.

• Side loads.

• Vertical loads. The front mount is secured to the intermediate case in three positions;

A monoball type universal joint. This gives the main support 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 gives resistance 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 the engine side to side movement and give vertical support.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Propulsion System Maintenance The following subjects are discussed in this section; Intake cowl Removal. AMM ref. 71-11-11-000-010. Installation. AMM ref. 71-11-11-400-010. Fan cowl doors Removal. AMM ref. 71-13-11-000-010. Installation. AMM ref. 71-13-11-400-010. Thrust reverser C ducts Removal. AMM ref. 78-32-01-000-010-left hand. AMM ref. 78-32-01-000-010-right hand. Installation. AMM ref. 78-32-01-400-010-left hand. AMM ref. 78-32-01-400-010-right hand.

Common nozzle assembly Removal. AMM ref. 78-11-11-000-010. Installation. AMM ref. 78-11-11-400-010. Note: Observe all safety precautions quoted in the AMM.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Inlet Cowl Removal and Installation AMM Ref. 71-11-11-000-010 The procedure to remove and install the inlet cowl is as follows;

• Open the L and R fan cowl doors.

• Attach the sling to the inlet cowl and the hoist.

• Remove the coupling at the anti ice duct joint and discard the seal. Fit new seal on installation.

• Disconnect the four electrical connectors at the top RH side of the cowl aft bulkhead.

• Disconnect the P2 signal pipe.

• Take the weight of the cowl on the sling with the hoist.

• Remove the cowl securing bolts.

• Move cowl forward carefully and lower onto dolly. Installation This is a reversal of the removal procedure. When offering up the inlet cowl use the 4 location spigots to ensure correct alignment. The following are the required test after installation; Engine air intake ice protection operational test. AMM ref. 30-21-00-710-001. P2/T2 operational test. AMM ref. 73-22-11-710-040.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Fan Cowl Doors Maintenance Warning Make sure that the landing gear ground safeties and the wheel chocks are in position. Be careful when opening the fan cowl doors in wind speeds of more than 30 mph but less than 60 mph. Injury to personnel and/or damage to the engine can occur. Do not open or allow to remain open fan cowl doors in wind speeds in excess of 60 mph. Injury and/or damage to the engine can occur. Fan Cowl Doors Opening AMM ref. 71-13-00-010-010 • Carry out the flight deck checks as per aircraft

preparation. • Ensure that the area around the engine is clear of

obstacles. • Open the latches starting from the front to the rear. • Engage the support struts to hold the fan cowl doors in

the open position. • Ensure that the support strut locking mechanisms are

secured. Fan Cowl Doors Closing AMM ref. 71-13-00-410-010 Hold fan cowl door to allow the disengagement of the support struts. • Lower the fan cowl door and align the locating pins.

• Fan cowl doors modified to SBN 71-0259 an additional

feature called the hold open device is fitted. To allow the fan cowl doors to come together fully depress the pin inwards on this device. This will allow the fan cowl doors to close.

• Engage the latches and close them in sequence from the rear to the front.

• Ensure that the fan cowl doors are located properly against the fan casing.

• Ensure that the closing forces exerted on the latches are within acceptable limits.

Note: There have been several instances over recent years, of aircraft experiencing Fan Cowl loss during take-off. This extremely hazardous situation has been the result incorrect maintenance practices. All instances of Fan Cowl loss have occurred on first flight after maintenance activity had recently taken place. SBN 71-0259 introduces a modification that is designed to make the fan cowl doors more prominent to the naked eye when they are open and in the down position. The fan cowl doors have a modification that gives them an open appearance when they are not closed and secured for flight.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Fan Cowl Doors Removal and Installation Removal AMM ref. 71-13-11-000-010 The procedure is summarised below.

• Remove the blanking caps from the cowl slinging points.

• Attach sling to cowl door and hoist.

• Open cowl door to gain access to hinges.

• Remove split pins from hinge bolts.

• Remove nuts and shouldered bolts.

• Remove cowl door and lower onto dolly. Installation AMM Ref. 71-13-11-400-010 This is the reversal of the removal sequence. On completion, check the cowl door alignment and latch tension. Note: The Fan Cowl doors weigh 93 to 105 lbs.(42kg to 47kg) If there is a strake fitted ref to AMM 71-13-19-000-010-A for removal/installation.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Thrust Reverser C-Ducts Maintenance Warning: The opening and closing procedure for the thrust reverser C-ducts must be adhered to fully. These units can close very quickly and neglect can cause injury to personnel. Opening AMM ref. 78-32-00-010-010 • Carry out the flight deck checks as per aircraft

preparation. • Ensure that the area around the engine is clear of

obstacles. • Ensure aircraft leading edge slats are retracted. • Remove the HAP at the top of the translating cowl if

the thrust reverser is in the deploy position. • Open the fan cowl doors (71-13-00-010-010). • Deactivate the HCU (78-30-00-040-012). • Open the latch access panel and engage the auxiliary

latch and take up the tension of the two thrust reverser halves.

• Release the latches in the following sequence; 3, 2 ,5, 4, 1.

• Dis-engage the auxiliary latch. • Attach the hand pump and extend the thrust reverser

C-ducts to the open position. • Engage the rear then the front support struts in position

and then decay the hydraulic pressure to rest the units on the support struts.

• Disconnect the hydraulic hand pump. Closing AMM ref. 78-32-00-410-010 • Carry out the flight deck checks as per aircraft

preparation.

• Engage the hand pump and open the thrust reverser C -ducts.

• Disengage the support struts and stow them.

• Allow the thrust reverser units to close. Note: The forward most latch must be in the locked position before closing.

• Engage the auxiliary latch assembly and draw the thrust reverser units together.

• Check front latch has not fouled.

• Disengage the hand pump and engage all latches and lock them in the following sequence; 1, 4, 5, 2, 3.

• Ensure latch unlock indicators are engaged.

• Disconnect auxiliary latch and stow.

• Close the thrust reverser access panel.

• Reactivate the HCU (78-30-00-010-010)

• Close the fan cowl doors (71-13-00-410-010).

• Return the aircraft back to its usual condition.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System C-Duct Maintenance Slinging and Hoisting After removal the C-ducts are mounted on to the transportation and work stand. IAE 1N20005 L/H and IAE 1N20006 R/H. Each C-duct is attached to the aircraft pylon by four hinges. The three front attachment points are provided by beams located on the bottom of the pylon. The beams are not rigidly attached to the pylon and this provides a degree of self alignment when closing the C-ducts. The rear hinge point is a solid location on the side of the pylon. Note: The hinged access panel must be removed to gain access to the thrust reverser C-duct hinges. The translating cowl must be in the stow position.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Latch Adjustment and Alignment Purpose Latch adjustment is carried out to ensure that the correct gap between fan cowl doors and thrust reverser C-ducts are achieved. The latches are set to achieve the desired clamping force required to satisfactorily hold the fan cowl doors and thrust reverser C-ducts closed. Location The latches are located at bottom dead centre (BDC) of the fan cowl doors and thrust reverser C-ducts. Fan Cowl Doors Fan Cowl latch adjustment for ‘into’ and ‘out of’ wind step is carried out by adjusting the nuts that attach the latch keeper to the keeper housing. “into” and “out of” wind checks ref to AMM 71-13-00-991-155. > than 0.040 in (1,02mm) out or > than 0.050 in (1,27 mm) in adjust latches ref to AMM 71-13-00-800-012. Latch tension is adjusted by use of the adjusting nut at the back of the latch keeper. The latch closing load should be between 45 to 55 lb. (20.02 daN – 24.47 daN). Thrust reverser C-Ducts Thrust reverser C-ducts latch adjustment for into and out of wind step is carried out by adjusting the nuts that attach the latch keeper to the keeper housing. Latch tension is adjusted by use of the adjusting nut at the

back of the latch keeper. The latch-closing load should be between 45 to 55 lbf. (20.02 daN – 24.47 daN).




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Combined Nozzle Assembly (CNA) Removal AMM Ref. 78-11-11-000-010

• Lift the IAE 1N20001 CNA Fixture up to the CNA and secure with straps.

• Disconnect the ACAC exhaust duct.

• Support the weight of the CNA (approximately 213 lbs.) and remove the 56 nuts and bolts.

• Lower the CNA fixture onto the IAE 1N20004 CNA dolly. Installation AMM Ref. 78-11-11-400-010 Refitting is the reverse of the above steps.

Revision 2 Page2-33


Revision 2 Page2-34

© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Engine Combined Drains System Purpose To provide an early indication of a system or component 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. It protrudes from the bottom of the fan cowl doors. Description This provides a combined overboard drain through a drains mast. The drains are for fuel and oil from the core module components, the LP compressor/intermediate case components and the external gearbox.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Engine Drains System Schematic The engine drains system schematic is shown on next page. For the accept/reject standards consultation of the AMM is recommended. For information and training reference only an extract of the AMM is provided below.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Propulsion System Engine Storage Caution: You must keep the engine in storage for too long.

The times given in the procedure are the maximum for which the engine can be preserved. If the time engine is in preservation is to be extended, you must do the full preservation procedure again. If these procedures are not followed damage to the engine can occur.

Caution: You must do all the applicable procedures when an

engine is put into storage. If they are not, corrosion and general deterioration of the core engine and the fuel system can occur.

The task 72-00-00-500-001 gives details of the required procedures for preservation and storage of the engine or QEC unit that is to be stored or transported. Protective treatment for the engine is dependant on the climatic conditions in which the engine is to be stored. Refer to task 72-00-00-500-002.

Prepare the engine for storage. Additional storage requirements refer to fig below. 72-00-00-990-243. Note:

1. The use of VMI bags affords maximum protection to the engine/QEC unit and must be utilised wherever possible, regardless of the storage environment and time period.

2. Use a full polythene cover or similar, secured around the engine, and engine stand preventing the ingress of dirt, grit and sand.

3. If the same conditions can be achieved, without the use of a VMI, use full engine protection from direct and indirect moisture as well as protection from adverse weather conditions and ingress of any type, then this is allowed.

Desiccant must still be used in accordance with TASK 72-00-00-500-005 and the integrity of the engine covers must be checked periodically






Cooling & Ventilation (Chapter 75)

Fire Protection (Chapter 26)

© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Fire Protection and Ventilation The purpose of fire protection is to give an indication to the flight deck of a possible fire condition about the engine. The purpose of the ventilation system is to provide a flow of cooling air about the engine to reduce the risk of a fire condition annunciation to the flight deck. Location The locations of the fire detection fire wires are about the fan casing and core engine. The location of the ventilation air is about the entire of the fan case and core engine. Description The engine is ventilated to provide a cooling airflow for maintaining the engine components within an acceptable operating temperature. Also to provide a flow of air that assists in the removal of potential 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 exits at the exhaust located on the bottom rear centre line of the fan cowl doors. Zone 2 ventilation Metered holes within the inner barrel of the “C” duct allow pressurized fan air to enter the zone 2 area. Air exhausting from the active clearance control (ACC) system around the turbine area also provides ventilation air for Zone 2. The air circulates through the core compartment and exits through the lower bifurcation of the C ducts via the thrust recovery duct. Ventilation during ground running During ground running local pockets of natural convection exist providing some ventilation of the fan case zone 1. Zone 2 ventilation is provided by fan duct pressure as above, during ground running and flight.


© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Fire Protection and Ventilation


© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Fire Protection and Ventilation Fire Detection System Purpose The fire detection system monitors the air temperature in Zone 1 and Zone 2. When the air temperature increases to a pre determined level 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 combustor diffuser 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 triggering sensors. When the air temperature about the fan case and/or core engine increases to a pre-determined level the system is designed to detect this and display a warning message and 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 of each other. Each zone has two detector units which are mounted as a pair, each unit gives an output signal when a fire or overheat condition occurs. The two detector units are attached to support tubes by clips. Nacelle air temperature (NAC) Zone 2 has the nacelle air temperature sensor. Indication is to the flight deck when a temperature exceedance has occurred.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Fire Protection and Ventilation Fire Detection System and Detector Units The fire detection system employs detector units called firewires. The firewires are mounted in pairs. This is necessary due to the level 3 class 1 message that they generate when a fire 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 Warning

Computer (FWC). • The FWC generates the flight deck indications for a fire

condition. There is one FDU per engine. The FDU has two channels, each channel is looking at a separate fire detector loop of zones 1 and 2. Under normal conditions both firewires require to be indicating to the FDU to give a real indication to the flight deck. If there is a single loop failure of more than 16 seconds then the remaining firewire will continue to operate. The FDU will recognise the faulty fire loop. The faulty loop will be indicated to ECAM as the following message; ENG 1 (2) FIRE LOOP A (B) FAULT If there is a double loop failure then the FDU will recognise this as a possible burn through and the fire message will be generated to the flight deck.

Firewire detectors Each of the firewire detector units comprises of the following;

• A hollow sensor tube.

• A responder assembly. Sensor tube The sensor tube is closed and sealed at one end and the other open end is connected to the responder. The tube is filled with helium gas and carries a central core of ceramic material impregnated with hydrogen. An increase in the air temperature around the sensor tube causes the helium to expand and increase until the pressure causes the alarm switch to close. The FDU recognises this as an abnormal situation, hence fire indication will be illuminated. If a ‘burn through’ occurs, the pressure within the sensing tube is lost and as a result of this the integrity switch opens to give an indication to the FDU of a loop failure. Responder The responder has two pressure switches, one normally open and the other normally closed.

• The normally open switch is the alarm indication.

• The normally closed switch is the fault indication.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Fire Protection and Ventilation Fire Detection System Fire Bottles Purpose The fire bottles provide a means of extinguishing a potentially hazardous fire about the engine when a fire annunciation 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 found on 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 when the pressure falls below 225 psi. The indication in the flight deck is; AGENT 1 (2) SQUIB DISC This is an illuminating annunciator light on the overhead panel.

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

• The squib is activated from the flight deck.

• Excessive pressure in the fire bottle. 1600 to 1800 psi at 95 deg.C

The squib is an Electro Pyrotechnic Cartridge containing explosive powder. Two filaments ignite the powder when they are supplied with 28v dc. There is facility to carry out a fire system test that will give all the expected indications if all is functioning correctly. The fire test switch is located on the fire push button panel on the overhead panel.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Fire Protection and Ventilation Fire Detection System Indications and Controls Purpose The purpose of the fire detection system indications is to alert the flight crew to a possible fire condition. The controls allow the flight crew to react and deal with the impending fire indication in the flight deck. Location The fire control panel is located on the overhead panel for fire bottle operation and fire system test. The fire indication is located: Discreet warning light on the engine control panel located on the centre control pedestal. This is accompanied by other flight deck indications. Description The indication to the flight deck of an engine fire is a red warning. This level of alert is of the highest priority and requires immediate action. Engine fire warning When a fire or overheat is detected the following will occur in the flight deck; • Master switch light illuminates. • Fire discrete light illuminates. • Repetitive audible chime. • Engine fire push button illuminates. • ECAM warning message in red.

Reaction to fire warning The flight crew and ground test crews will react to the fire message by doing the following; • Depress the master warning light to silence the audible

chime. • Retard engine throttles to idle if power condition is

above idle. • Master lever set to off. • Select the engine fire push button to the out position.

By doing this the caution audible single chime alert will happen and the squib light will illuminate

• Wait 10 seconds to allow the engine to reduce in RPM. This will increase the extinguishing agent effect

• Discharge agent 1 and observe for agent 1 discharge light.

• Wait 30 seconds, if fire condition still exist then discharge agent 2 and observe for discharge light.

Note: Setting the push button to the out position will isolate the engine’s fuel, hydraulic, pneumatic and electrical power supplies from the aircraft. Fire warnings in flight and on the ground are the same.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Fire Protection and Ventilation Nacelle Air Temperature (NAC) Purpose The nacelle air temperature gives an advisory indication to the lower ECAM CRT if a temperature exceedance has been experienced. Location The NAC sensor is located by the bifurcation panel at bottom dead centre between the two thrust reverser C duct halves. The NAC is in zone 2. Description Under normal conditions the NAC indication is not displayed on the lower ECAM CRT. When a temperature exceedance of 320 deg.c has occurred the indication will appear to the lower ECAM CRT. This indication is displayed if; The system is not in engine starting mode and one of the two temperatures reaches the advisory threshold.




© IAE International Aero Engines AG IAE V2500 Line and Base Maintenance Fire Protection and Ventilation Nacelle Temperature Sensing and Fire Detection HarnessThe nacelle temperature sensing and fire detection harness electrical connections are shown below.




SECTION 3

MECHANICAL ARRANGEMENT

(Chapter 72)

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement

•

•

•

•

•

•

•

•

Mechanical Arrangement General The engine is an axial flow, high by-pass ratio, and twin spool turbo fan. The general arrangement is shown below. L.P. System L.P. compressor - comprising:

1 Fan stage L.P. Compressor (booster) consisting of (4 stages A5 derivative) (3 stages A1 derivative) driven by: Five stage L.P. Turbine Handling bleed valve at stage 2.5.

H.P. System • Ten-stage axial flow compressor driven by a 2 stage

H.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 combustion chamber, with 20 fuel atomizer type 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 the engine driven accessories and the pneumatic starter motor.

Revision 2 Page 3-1


Revision 2 Page 3-2


•

•

•

•

•

•

•

•

•

•

•

•

•

•

Engine Main Bearings The main bearing arrangement and the bearing numbering system is shown below. The 5 bearings are located in 3 bearing compartments:

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

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

The Rear Bearing Compartment located in the Turbine Exhaust Case houses the 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" oil damping.

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 a series of rod springs (squirrel cage).

No 4 Bearing • 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.

Revision 2 Page 3-3


Revision 2 Page 3-4

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement Bearing Compartments Front Compartment The No’s 1, 2 and 3 bearings are located in the front bearing compartment which is at the centre of the intermediate module (32). The compartment is sealed using air supported carbon seals, plus an oil filled (Hydraulic) seal between the H.P. and L.P. shafts. The 8th stage compressor air supports this seal. Adequate pressure drops across the seals to ensure satisfactory sealing are achieved by venting the compartment, by an external tube to the de-oiler. Gearbox Drive The HP Stubshaft, which is located axially by the Number 3 Bearing, has at it’s front end a bevel drive gear which, through the ‘Tower Shaft’ provides the drive for the Main Accessory Gearbox. The HP Stubshaft separates from the HP Compressor Module at the ‘Curvic Coupling’ and remains as part of the Intermediate Module.

Revision 2 Page 3-5


Revision 2 Page 3-6

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement Bearing Compartments Front Compartment (Continued) The drawing below shows details of the Number 2 and Number 3 Bearings. A Phonic Wheel is fitted to the LP Stub Shaft; this interacts with speed probes to provide LP Shaft speed signals (N1) to the Engine Electronic Control (EEC) (see section 11 – Engine Indicating). A speed signal is also provided to the Engine Vibration Monitoring Unit (EVMU), which is located in the Aircraft Avionics Compartment. The Hydraulic Seal prevents oil leakage from the compartment passing rearwards between the H.P. and L.P. shafts. The Number 3 Bearing is hydraulically damped. The outer race is supported by a series of eighteen spring rods, which allow some slight radial movement of the bearing. The bearing is centralised by rods and any radial movement is dampened by oil pressure fed to an annulus around the bearing outer race. The gearbox gear is splined onto the H.P. shaft and retained by the Number 3 Bearing Nut.

Revision 2 Page 3-7


Revision 2 Page 3-8

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement No 4 (Centre) Bearing Compartment The No 4 bearing compartment is situated in an inherently hostile, high temperature and pressure environment at the centre of the combustion section. The bearing compartment is shielded from radiated heat by a heat shield and an insulating supply of relatively cool air. This supply of cooled air (called 'buffer air') is admitted to the space between the chamber and first heat shield. The buffer air is exhausted from the cooling spaces close to the upstream side of the carbon seals, creating an area of cooler air from which the sealing function is obtained. This results in an acceptable temperature of the air flowing across the face of the carbon seals into the bearing compartment. Restrictors at the outlet from the cooling passage control buffer airflow rates. The bearing compartment internal pressure level is determined by the area of the variable scavenge valve. (No 4 Bearing Scavenge Valve described in the oil system). Essentially this valve acts as a variable flow restrictor in the No 4 Bearing Compartment vent line.

Revision 2 Page 3-9



© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement No 5 Bearing (Rear) Bearing Compartment The rear bearing compartment is located at the centre of the L.P. turbine module (module 50) and houses the No 5 bearing which supports the L.P. turbine rotor. An air supported (Stage 8) carbon seal seals the compartment at the front end. At the rear is a simple cover plate, with an ‘O’ ring type seal, secured by twelve bolts. Inside the compartment sealing is achieved on the LP shaft end by a small disc type plug, with a ‘spring supported jacket cup’ ring seal secured by a double helix spring clip. There are no air or oil flows down the LP shaft. Separate venting is not necessary for this compartment because with only one carbon seal, the airflow induced by the scavenge pump provides the required pressure drop across the seal. The pressure supply and scavenge oil pipes are covered by an insulating heat shield material.





•

•

•

•

•

•

•

•

•

•

•

•

•

•

•

•

Engine Internal Cooling and Sealing Airflows Purpose Sealing airflows provide positive air pressure to the bearing chambers to prevent oil loss. Cooling airflows provide cooling air for the engines internal components keeping them within designed operating temperatures. Location The air used for internal cooling and sealing is taken from the compressor stages of:

LPC stage 2.5. HPC stage 6 (A1only). 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 turbine blades. Air through the Air Cooled Air Cooler (ACAC). This is used for the cooling of the ‘buffer air’.

Buffer air is used to provide: Cooling, sealing and scavenge air for the No.4 Bearing Chamber.

LPC stage 2.5 air is used to: • Seal for the front and rear of the Front Bearing

Chamber. Note: (HPC stage 6 air seals the FBC on the early A1 engines only).

HPC stage 7 air is used for: • Airflow control for compressor stability, thermal anti-

icing and aircraft services bleed supply. HPC stage 8 air is used to:

Seal between the LP & HP shaft in the Front Bearing Chamber at the hydraulic seal and the sealing at the front of No. 5 Bearing Chamber.

HPC stage 10 air is used for:

Airflow control and aircraft services supply. ‘Make up’ air supply for the HPT stage 2 disc and blades. 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 and sealing of the no. 4 bearing chamber.





•

•

•

•

•

•

•

•

•

•

Modular Construction • 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 Designation Module No Module 31 - Fan 32 - Intermediate 40 - HP System 41 - HP Compressor 45 - HP Turbine 50 - LP Turbine 60 - External gearbox

Note: The module numbers refer to the ATA chapter reference for that module. 40 HP System • 41 - HP Compressor. • 42 - Diffuser Case and Outer combustion liner. • 43 - No 4 Bearing. • 44 - Stage1 Turbine Nozzle Assembly. • 45 - HP Turbine





•

•

•

•

Module 31 Description Module 31 (Fan Module) is the complete Fan assembly and 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 dovetail root. The front and rear blade retaining rings provides axial retention. Removing the front blade retaining ring and sliding the blade along the dovetail slot in the disc easily achieve blade removal/replacement. 22 annulus fillers form the fan inner annulus. The nose cone and fairing smooth the airflow into the fan.





•

•

•

•

•

•

•

•

•

•

•

•

Module 32 - Intermediate Case The Intermediate Module comprises of:

Fan Case

Fan Duct

Fan Outlet Guide Vanes (OGV)

LP Compressor (A5 variant - 4 stages)

(A1variant – 3 stages)

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 of

the 'C' ducts





•

•

•

Module 32 - Intermediate Case Instrumentation

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

P12.5

P2.5

T2.5

The rear view of the intermediate case is shown below.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement Module 40 HP Compressor Description

The HP compressor assembly (Module 40 is a 10 stage axial flow compressor. It has a rotor assembly and stator case. The compressor stages are numbered from the front, with the first stage is stage being designated as stage 3 of the whole engines compressor system. Airflow through the compressor is controlled by variable inlet guide vanes (VIGV); variable stator vanes (VSV) and handling bleed 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 the rotor. The compressor blades in stages 3 to 5 are attached to the compressor disks in axial dovetail slots and secured by lockplates. The stages 6 to 12 compressor blades are installed in slots around the circumference of the disks through an axial loading slot. Lock blades, lock nuts and lock screws hold the blades in position. The HP compressor stator case has two primary sub-assemblies, the HP compressor front and rear cases. Revision 2 Page 3-23



© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement Module 40 HP Compressor The HP compressor front case assembly has two split cases bolted together along the engine horizontal centre line. The front case assembly contains the VIGV’s, the stages 3 to 5 VSV’s and the stage 6 stator vanes. The front lower outer case provides a mounting for the VIGV and VSV 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 ring cases and an outer case. Flanges on the inner cases form annular manifolds, which provide stages 7 and 10 air offtakes. The five inner cases are bolted together, with the front support cone bolted at the stage 7 case and the stage 11 case bolted to the rear outer case. The five inner cases contain the stages 7 to 11 fixed stator vanes. The rear outer case is bolted to the diffuser case and to the rear flange of the HP compressor front case. Access is provided in the compressor cases for borescope inspection of the compressor blades and stator vanes





•

•

HP Compressor Compressor Drums - (Rotor) The rotor assembly is in two parts-

The stage 3 to 8 drum

The stage 9 to 12 drum

The two rotor drums are bolted together with a vortex reducer installed between the 8 and 9 stages. The vortex reducer straightens the stage 8 airflow, which passes to the centre of the engine for internal cooling and sealing.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement HP Compressor-Blades The compressor blades in stages 3 to 5 are attached to the discs in axial dovetail slots and secured by lock plates. Rubber strips bonded to the underside of the platform seal gaps between the blades. The stages 6 to 12 are installed in a slot around the circumference of the discs. Each disc has one axial loading slot to enable the blades to be installed into the disc. Four lock blades are installed on each disc, two on each side of the loading slot, which are locked by lock nuts and jackscrews.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement Combustion Section The combustion section includes the diffuser section, the combustion inner and outer liners, and the No 4 bearing assembly. Diffuser Casing The diffuser section is the primary structural part of the combustion section. The diffuser section has 20 mounting pads for the installation 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 pass through the diffuser casing. The inner combustion liner is attached to the turbine nozzle guide vane assembly. The inner and outer liners are manufactured from sheet metal with 100 separate liner segments attached to the inner surface (50 per inner and outer liner). The segments can be replaced independently during engine overhaul.





•

•

•

Turbine Nozzle Assembly The drawing below shows the arrangement of the diffuser casing and the inner and outer combustion liners, the No1 NGV’s, and the TOBI (Tangential Out Board Injector) ducts. Also shown is the No 4 bearing support assembly. The primary parts of the Stage 1 Turbine Nozzle Assembly

The Stage 1 HPT Vane Cluster Assemblies

The Stage 1 HPT Cooling Duct Assembly

The Combustion Chamber Inner Liner

The stage 1 turbine nozzle assembly has 40 air-cooled vanes, made of cobalt alloy. The vanes are attached to the stage 1 HPT cooling duct assembly with bolts. The stage 1 has 40 vanes; each hollow vane has internal baffles and cooling holes in the airfoil. Vane airfoils also have a heat-resistant coating. The stage 1 vanes are held in position by the stage 1 HPT cooling duct assembly. The duct is installed on the rear-inner flange of the diffuser case. Operation The ring of vanes makes a series of nozzles, which increases the velocity of the gases from the combustion chamber. The vanes direct the combustion chamber

gases at the optimum angle onto the stage 1 turbine blades. The internal vane baffles and airfoil cooling holes permit relatively cool air from the diffuser case to go through the vane and over the external airfoil to decrease metal temperature. Sheet-metal seals between adjacent vane platforms decrease leakage of the cool air.





•

•

•

•

•

•

•

•

•

HP Turbine Description

The primary parts of the HP turbine rotor and stator assembly are:

The HP Turbine Rotor Assemblies (Stage 1 and 2)

The HP Turbine Case and Vane Assembly The HP turbine rotor assemblies are two stages of turbine hubs with single-crystal, nickel-alloy blades. The two-hub configuration removes a bolt flange between hubs. This decreases the weight and enables faster engine assembly. The blades have airfoils with high strength and resistance to creep. Satisfactory blade tip clearances are supplied by Active Clearance Control (ACC) to cool the case with compressor air. The primary parts of the stage 1 rotor assembly are:

Stage 1 Turbine Hub Inner and Outer HPT Air Seals 64 Blades Rear HPT Air Seal

The primary parts of the stage 2 rotor assembly are:

Stage 2 Turbine Hub 72 Blades Stage 2 Blade Retaining Plate

The inner and outer HPT air seals are installed on the front of the stage 1 hub. The stage 1 blades are installed in slots on the hub. The blades are held on the forward side by the outer HPT air seal. The stage 2 HPT air seal is installed on the rear of the stage 1 hub. This air seal holds the stage 1 blades on the rear side. Stage 1 blades and Nozzle Guide Vanes are cooled using H.P. Compressor discharge air. The stage 2 turbine hub is installed behind the stage 1 hub and the stage 2 HPT air seal. Stage 2 blades are installed in slots in the hub. The blades are held on the forward side by the stage 2 HPT air seal. The blades are held on the rear side by the stage 2 blade retaining plate. Stage 2 HPT blade cooling air is a mixture of HPC discharge air and stage 10 compressor air. This air passes through holes in the stage 1 HPT (front inner) air seal and the stage 1 turbine hub into the area between the hubs. The air then goes into the stage 2 blade root and out the trailing-edge cooling holes.





•

•

•

•

•

•

•

•

•

LP Turbine Description

The primary parts of the Low Pressure Turbine (LPT) 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 power to the LP compressor through the LPT shaft. The LPT rotor 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 welded assembly. To identify the LP turbine module, an identification plate is attached to the LP turbine case at the 136degrees position. The LPT case has two borescope inspection ports at 125.27 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

The five LPT disks are made from high heat resistant nickel alloy. The LPT blades are also made from nickel alloy and are attached to the disks by firtree type roots. The blades are held in axial position on the disk by the rotating air seals (knife-edge).





•

•

•

•

•

•

•

•

•

•

•

•

Module 60 - External Gearbox Purpose The gearbox assembly transmits power from the engine to provide drives for the accessories mounted on the gearbox front and rear faces. During engine starting the gearbox also transmits power from the pneumatic starter motor to the core engine. The gearbox also provides a means of hand cranking the HP rotor for maintenance operations. Location The gearbox is mounted by 4 flexible links to the bottom of the fan case.

Main gearbox 3 links. Angle gearbox 1 link.

Description The external gearbox is a cast aluminium housing that has the following features;

Individually replaceable drive units. Magnetic chip detectors. Main gearbox 2 magnetic chip detectors. Angle gearbox 1 magnetic chip detector.

The following accessory units are located on the external 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 is provided by a combination of carbon and ‘O’-ring type seals. The carbon seals can be replaced while the engine is on wing.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement Engine View Right Hand Side The following components are located on the right hand side of the engine. 1. Stage 10 make-up air valve for supplementary turbine

cooling. 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. 10th stage 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.

19. LPT and HPT active clearance control valves (ACC). 20. HPC stage 10 handling bleed valve. 21. Engine rear mount. 22. Booster bleed valve slave actuator. 23. Front engine mount. 24. HPC 10th stage cooling air for the HPT 2nd stage NGVs. 25. Solenoids for the three off HPC 7th stage handling

bleed valves. 26. Solenoid for the HP10 make-up cooling air control

valve. 27. Solenoid for the HP10 cabin bleed PRV/Shut-off valve.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement Engine View Left Hand Side The following components are located on the left-hand side of the engine. 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 10th stage cabin bleed offtake pipe. 19. HPC 10th stage pressure regulating/shut-off valve

(PRSOV).

20. HPC 7th stage bleed valve (HPC7 C). 21. HPC 7th stage cabin bleed non-return valve (NRV). 22. VIGV/VSV actuator. 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.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement Borescope Plug Access The borescope plugs for the compressors; combustor and turbines are mainly found on the right hand side of the core engine. The exception being the combustor and turbines, these access positions are found on both sides of the core engine. LP Compressor Borescope Access A1 engines Borescope access is possible for stages 1.5 and 2.5 only. There are no access features to remove. Guide tubes and fibrescopes are used for the inspection. A5 engines Borescope access is possible for all stages of the LPC booster. There is one access port that requires the removal of two FEGVs clusters at approximately 5 o’clock position, when viewed from the rear. This will give access to the trailing edge of stage 2.0 and the leading edge of stage 2.3.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement A5 Engines HP Compressor Borescope Access There are nine borescope access ports for the HP compressor. Three of these are located about access position B. HPC stage 3F Port A.

HPC stage 3R and 4F Port B.

HPC stage 5R and 6F Port C.

HPC stage 7R and 8F Port D.

HPC stage 8R and 9F Port E.

HPC stage 9R and 10F Port F

HPC stage 11R and 12F Port G.

Where ‘F’ denotes the front of that particular stage. Where ‘R’ denotes the rear of that particular stage. Note: During the removal of the borescope ports the old jointing compound must be cleaned off. Before installation of the borescope ports jointing compound must be used as recommended by the AMM. Take care not to let excessive jointing compound enter the borescope access port hence into the engine




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Mechanical Arrangement Combustor, HP and LP Turbines Borescope Access Borescope access for the combustor is found in eight positions, of which six are found around the combustion outer case and the addition of the two igniter ports. Combustor SB 72-0221 introduces a new diffuser case assembly. A1 Diffuser Case (Pre SB 72-0221) Access to inspect the combustion chamber and the HPT stage 1 vanes is by 5 plugs with gaskets. These are numbered:

• B1 to B4 for the left hand side of the engine.

• B5 and the 2 igniter plug ports for the right hand side of the engine. A1 Diffuser Case (Post SB 72-0221) Access to inspect the combustion chamber and the HPT stage 1 vanes is by 6 plugs with gaskets. These are numbered:

• B1 to B5 for the left hand side of the engine.

• B6 and the 2 igniter plug ports for the right hand side of the engine. Note: The borescope access ports are located near the diffuser case rear flange. The ports must not be confused with the 5 larger locating pins that are equi spaced around the forward end of the case.

HP Turbine The HP turbine has provision for inspection of the leading and trailing edges of the blades. LP Turbine The LP turbine has borescope inspection for the stage three leading edge only. Note: When installing borescope access features to the combustion system and HPT stage 1 the threads of the fasteners must be coated with an anti galling compound and an anti seizure compound as recommended by the AMM. When installing borescope access features to the HPT stage 2 and LPT stage 3 the threads of the fasteners must be coated with engine oil as recommended by the AMM.




SECTION 4

LP COMPRESSOR (FAN) MAINTENANCE

(Chapter 71)

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Nose Cone The Glass-fibre cone smoothes the airflow into the fan. It is secured to the front blade-retaining ring by 24 bolts. A Fairing is attached to the front blade-retaining ring by 6 bolts. Note: Balance weights must not be placed at these 6 bolt locations on the fairing. The Nose Cone is balanced during manufacture by applying weights to its inside surface. The nose cone is un-heated. A soft rubber cone tip provides ice protection. As ice builds up on the tip, it becomes un-balanced and flexes. This causes the ice to be dislodged from the rubber tip and is then ingested by the fan before it has built up to a significant mass. The Nose Cone retaining bolt flange is faired by a titanium fairing which is secured by six bolts. The arrangement is shown below. Note: Take care when removing the Nose Cone retaining bolts. Balance weights may be fitted to some of the bolts. The position of these bolts with their respective weights must be marked before removal, so as to ensure they are refitted to the same position.

Revision 2 Page 4-1

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance

Revision 2 Page 4-2

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Front Blade Retaining Ring The Assembly is shown below. The Front Blade Retaining Ring is secured to the Fan Disk by a ring of 36 bolts. A second (outer ring) passes through the retaining ring and permits the individual securing of the Annulus Fillers by 22 bolts. Both these sets of bolts must be removed before attempting to remove the Front Blade Retaining Ring. After the removal of the 22 annulus filler securing bolts and all 36 retaining ring bolts, it is possible to remove the front blade retaining ring by the use of 6 ‘pusher bolts being inserted into 6 threaded holes designed specifically for this purpose. Note: The fan blades and annulus filler positions are not identified. For this reason it is important to identify and make a note of the original blade and annulus filler positions prior to their removal. When the Nose Cone is fitted, it is possible to identify the positions of blades numbers 1,2 and 3 by noting that the front blade retaining ring has etched on it’s outer edge these blade number positions. These numbers are marked in a counter-clockwise direction when viewing the engine from the front. Having established the original positions of the blades it is important to number the blades and their corresponding annulus filler by using an approved marker pen (Material VS 06-69 ref 70-30-00). Revision 2 Page 4-3


Revision 2 Page 4-4

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Fan Trim Balance Procedure Reference 71-00-00-860-010 Engine Operation Limits, Guidelines and Special Procedures. Vibration limits and fan trim balance vibration guidelines: a) Vibration Limits (Steady State)

− N1 (peak): 5.0 Units.

− N2 (peak): 5.0 Units. 1. Engines that have vibrations within the vibration limits

are acceptable. 2. Engines that have N1 peak vibrations that exceeds the

above limit, troubleshoot as per Trouble Shooting Manual (ref. TSM task 77-30-00-810-826) or (ref. TSM task 77-30-00-810-827).

3. Engines that have N2 peak vibrations that exceeds the above limit, troubleshoot as per Trouble Shooting Manual (ref. TSM task 77-30-00-810-828) or (ref. TSM task 77-30-00-810-829).

4. A non-revenue ferry flight to a maintenance base is permissible with N1 or N2 vibration above limits, if no fault in the respective trouble shooting procedures in steps 2 and 3 above. This condition is permissible for only one engine per aircraft.

b) Vibration guidelines

− N1 (peak): 2.0 Units 1. Fan trim balance is recommended any time N1 peak

vibration exceeds this 2.0 Unit guideline. Perceivable airframe vibrations generally accompany N1 vibration levels above this guideline value. Waiting until N1 peak vibration approaches or exceeds the 5.0 unit limit may require multiple fan trim balances to bring N1 vibration down to an acceptable value.

Note: 5.0 units (Aircraft ECAM display) = 1.5 inches per second of displacement due to the imbalance. 2. Aircraft/Flight Crew Operating Manual (FCOM)

correlation.

− As stated in the FCOM, if N2 vibration during engine start exceeds limit, the start should be aborted. Subsequent starts may be initiated without maintenance action for up to three start attempts.

− The above limits and guidelines are stable (steady state) and as such may not be stable, therefore the aircraft level is advisory and not a limit. Vibration above the advisory level may or may not require maintenance action, as described in the FCOM; initially depending on icing conditions or other engine parameter shifts and finally if the advisory level is confirmed at steady state conditions.

Revision 2 Page 4-5


Revision 2 Page 4-6

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Fan Trim Balance There are two methods available to balance the fan, the ‘one shot’ and ‘trial weight’ the method. Both use data gained from the Engine Vibration Monitoring system (EVMS). The one shot method allows balancing of the fan with fewer engine ground runs required and has proved itself effective in service use. If necessary a Vibration survey (Test No 8) may be performed to obtain the vibration characteristics of the engine. Note: • If vibration exceeds limits during the survey ground run,

slowly bring engine speed to idle and shutdown.

• Angles are counter clockwise viewed from the front of the engine.

Data: (speed, amplitude and phase angle) may be collected on ground or during cruise flight, collection in flight is either automatic or for selected speeds and on the ground may be manually selected ref: AMM. Best results are obtained from data in the 80-90% N1 speed range with 85% N1 being the best single speed point, for ground running an average of correction.

Caution: Operate both engines for this test with the non-test engine set at 1.25 EPR for aircraft stability. Engine speed greater than 85% N1 (4645 rpm) can cause aircraft buffeting. An N1 Keep-Out-Zone (KOZ) of 61-74% N1 (AOW1056) has been introduced during all stages of engine operation on the ground, including Ground Testing, Taxiing and ‘Hold’ periods. This is to prevent the blade from experiencing high stresses as a result of ‘Fan Blade flutter’. This is particularly acute during ‘cross-wind’ conditions. The KOZ is being incorporated into EEC software (availability A5 SCN17-3rd Qtr 2002) For cruise flight, data at 5 speeds, pre-selected or automatically is collected. Data stored in the memory of the EVMU is accessed through the MCDU menu in the flight deck and should be printed for later reference and calculation.

Revision 2 Page 4-7


EVMU Engine Unbalance

Read Eng 2

N1 DISP PHASE DATE RPM MIL DEG D/M 3041 0.2 + 0 03/01 NO ACQUISITION 4199 0.5 +230 03/01 4524 0.5 +236 03/01 5088 0.6 +189 03/01

< RETURN

EVMU PRINTOUT FROM FLIGHT Revision 2 Page 4-8

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance One Shot Method The following procedure may be used to trim balance an engine fan whilst mounted on the aircraft wing. The data collection will be via the aircraft EVMU system. Data may be collected during a ground run or in cruise flight. Definitions • Speed (N1) expressed as a percentage 100% = 5650

rpm. Note! (1% N1 = 56.5 rpm)

• Amplitude (U) indicated vibration levels expressed in Mils (P-P) from the EVMU system.

• Phase Angle (A) indicated angle in degrees from the EVMU system.

• Phase Lag (B) dynamic phase lag of the LP system between phase angle and true position of unbalance.

Mass Coefficient (K) value by which the amplitude must be multiplied to give correction mass required or a given speed Fan Trim Balance with the EVMU (One Shot Method) Task (77-32-34-750-010) This procedure can be used for consecutive fan trim balances if necessary. If consecutive fan trim balances with this method do not give significant results, carryout a fan trim balance with the ‘Trial Weight’ method. Reference Task (77-32-34-750-010-01) to carryout either of these procedures, flight or ground vibration data must be available. Reference Task 77-32-34-869-048 (Unbalance data.

Acquire in flight, read on ground) or Task 77-32-34-869-010 (Acquisition of unbalance data on ground). Some aircraft are fitted with software (customer option) which permits the engineer to interrogate via the MCDU the stored data regarding out of balance correction required. This information is contained in the EVMU and by accessing the EVMU Engine Unbalance menu, it is possible to establish the necessary adjustments required to eliminate out of balance situations. Note: Prior to carrying out any adjustments, the engineer must first confirm the accuracy of the current status regarding the configuration of weights (position and part number) that are already installed and recorded in the system. To accomplish this it is necessary to physically verify the position and part number of the balance weights already installed onto the front blade-retaining ring.

Revision 2 Page 4-9








Annulus Fillers After removal of the Front Blade retaining ring 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 degrees counter-clockwise

• draw the Annulus Filler forward to clear the blades Remove the annulus fillers on either side of the blade to be removed. The blade to be removed can than be pulled forward to clear the dovetail slot in the fan disc. Examine the outer surface of the Annulus Filler for cracks, nicks, dents and scores. Limits in the AMM can be applied to assess the damage for accept or reject. If the surface coating of the annulus filler is damaged to the point of requiring a repair the AMM has a procedure that allows this to be done. AMM ref 72-31-11-300-010 gives comprehensive instructions as to the correct procedure for repair.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Annulus Fillers Caution: When re-fitting the Annulus Fillers, it is extremely important that correct location of the Annulus Fillers into the Rear Retaining Ring is achieved. If the Annulus Filler is not correctly installed, it is possible that when the Front Retaining Ring is subsequently torque tightened in place onto the Fan Disk, it may result in the deformation and displacement of the Rear Retaining Ring. This could cause it to come into contact with the inlet housing of LP Compressor Module.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Reposition of the Annulus Filler Seals AMM (72-31-11-300-017) During the installation of the Annulus Filler it is possible to cause the sealing strips to be incorrectly seated. If this were to be left uncorrected, it is possible that the Fan Blade would be displaced slightly prevented from it’s normal radial operating position. This in turn would cause the Fan Module to become un-balanced and vibration levels for the engine could be exceeded. The task referenced above documents the procedure to eliminate this. The task requires a stiff plastic strip to be used to reposition the seals if they ‘ rolled’ as shown in the diagram below. Note: An expired credit card is suitable, or a plastic checklist card Caution: Make sure the plastic strip has a smooth surface and edges. If you use a strip with a rough edge surface or edges, damage to the seal can occur. Make sure that you do not break the plastic strip and leave pieces of it in the Fan. Pieces of plastic can damage the rubber.

Procedure • Push the plastic between the Fan Blade and the

Annulus Filler at the rear of the Fan.

• Note: If this is difficult to do, it can be an indication that the seal is caught.

• Slide the plastic strip forward to move the seal into the correct position. Accomplish this procedure on both sides of the Fan Blade starting at the trailing edge of the blade and moving it forward to the leading edge.

• Note: When the seal is in the correct position you can easily move the plastic strip from the front to the rear of the blade.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Fan Blade Inspection Fan blade inspection procedures are briefly described in these notes. This information is for guidance only and the AMM ref Chapter 71-31-11-200-010 should be used as the reference document. General The fan blade surface area is divided into zones. The zones are; • Ar. • At. • Br. • Bt. • Cr. • Ct. • F. The acceptance limits for damage vary depending on which zone is damaged.

Inspection Standards Blades are inspected for signs of the following; • Nick’s. • Cracks. • Dents. • Scores. • Surface scratches. • Bends on the leading or trailing edges.

• Arc burns (lightning strikes). Any blade, which has Arc burns or cracks must be rejected and a replacement blade fitted. An Arc burn is evident by a small circular or semi-circular heat affected area of the blade surface that may contain a shallow pitting, remelting or cracking. Visually a dark blue discoloration is associated with the heat-affected area




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Fan Blade Inspection General The leading and trailing edges of the fan blades should be examined for bends (deformations).

• The maximum number of bent blades in a fan rotor assembly is three.

• No more than one bend in a blade is permitted.

• If any bend has associated cracks, kinks, creases tears or nick’s then the blade must be rejected as per AMM recommendations.

• Bends must be outboard of the annulus fillers, if any bend extends below the annulus filler platform, reject the blade as per AMM recommendations.

• Any blade untwist is acceptable as per AMM recommendations.

• No bending is acceptable in the area F as per AMM recommendations.

• There must be a smooth transition between the undamaged airfoil surface and the bent area. If there is not a smooth transition reject the blade as per AMM recommendations.

Acceptance Limits X maximum = 0.2 in. (5,08 mm) Y must not be less than 8 times dimension X if Y is less than 8 times X, reject the fan blade. Z must not be less than 15 times dimension X if Z is less than 15 times X, reject the fan blade.

Fly Back Limits Accept dimension X between 0.2 and 0.5 in. (12,7 mm) providing dimensions Y & Z follow the same criteria as above. The blade must be changed within 125 hours or 25 flights, whichever is the sooner as per AMM recommendations.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Introduction of LPC Rotor Balancing Procedure in Fan Blade Replacement. AMM 72-31-11-400-010 When a replacement blade is installed, procedures to keep the balance of the LP compressor rotor are necessary. The correct method used depends on the difference between the old and new blade “weight and moment”, which is etched onto the bottom of the fan blade root as shown below.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Introduction of LPC Rotor Balancing Procedure in Fan Blade Replacement. AMM 72-31-11-400-010 When a replacement blade is installed, procedures to keep the balance of the LP compressor rotor are necessary. There are four methods used to correct the balance in this task. These are:

• Method 1.

• Method 2.

• Method 3.

• Method 4. The methods can be selected according to the conditions of damage seen on the LPC rotor. Method 1 Method 1 uses the trim balance weights on the 36 bolt hole flange (front blade retaining ring) to compensate the moment weight difference between the removed and installed blade. Method 2 Method 2 uses the balance weights on the 22 bolt hole flange (front blade retaining ring) to compensate for the moment weight difference.

Method 3 Method 3 uses a pair of replacement blades. The moment weight difference of the damaged blade position is compensated by non-damaged blade replacement at a diametrically opposite position. Method 4 In method 4, the distribution of all the 22 fan blades is changed. Removal and installation of all blades is necessary. When two or more blades are replaced, select the applicable method in each case if method 4 is not used. Note: Seven methods of fan blade installation procedure are given in this task. Four of the seven methods are used when balance correction is applied. The remaining three methods are used when the balance correction is not applied. (see chart below)




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Fan An example of balance weight fitted to positions on the 22 Bolt Hole Flange. In this example the balance weights are positioned to supplement the radial moment weight of a replacement blade that has a lower radial moment weight than that of the damaged blade replaced.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Fan Blade Inspection - TAP Test The Transient Acoustic Propagation (TAP) test of the fan blades is detailed in Maintenance Manual Ch 72-00-00 Inspection/Check 09 PB600. TAP Tester System Check

• Connect probe to tester.

• Press the ON switch, the display shows RRM THOR UNIT

• Press the MENU switch, the display shows SYSTEM TEST.

• Press the EXEC switch, the display shows the SYSTEM OK.

• Press the OFF switch. Functional Check of the TAP Test Set • Apply a small amount of ultrasonic couplant to the TAP

test block.

• Put the test block on a flat surface and attach the probe to the centre of the test block.

• Press the ON switch, the display shows RRM THOR UNIT.

• Press the EXEC switch, the display shows a value - make sure this is within the values engraved on the side of the test block.

Inspection of the Fan Blades

• Apply a small amount of ultrasonic couplant to the lower convex airfoil adjacent to the annulus filler (as shown).

• Attach probe to the fan blade.

• Press the ON switch.

• Press the EXEC switch. The display will show the value or message after approximately 4 seconds.

If the display shows more than 700 dB/sec, reject the engine as per AMM recommendations.

• Repeat for all 22 fan blades.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Fan Blade Repairs Detailed information regarding the repair of damage on the Low Pressure Compressor (LPC) Fan Blades by local material removal can be found in the AMM Task (72-31-11-300-016). Repair Scheme VRS1506. Caution: • The maximum number of dressed blades for a

given Compressor Fan Blade set, is the equivalent of three blades dressed to the maximum limit. The remaining blades must not be dressed.

• Titanium component – You must use silicon carbide type abrasive wheel stones and papers to dress, blend and polish the Fan Blade.

• Titanium component – Do not use force with mechanical cutters or the material will become too hot.

• Titanium component – If the material shows a change in colour to darker than a light straw colour, the Fan Blade is to be rejected.

Note: The repair scheme VRS1506 allows scalloping of the leading edge of the fan blade. Remove damage from the airfoil surface and if damage is found in Zone ‘AD’ then you must blend parallel with the leading edge, by removing material above the repaired area.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Fan Blade Root Dry Film Lubricant Inspection AMM Task (72-31-11-200-012) Examine the Blade Root of the Stage 1 Fan Blade for Dry Film Lubricant peeling. If the dry film lubricant shows any sign of peeling, carryout a repair of the coating as per AMM recommendations.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Repair of Fan Blade Chocking Pads AMM (72-31-11-300-019)

Repair Scheme Number VRS 1063 Should the Chocking Pads become detached from the Fan Blade, it is possible to carryout a repair utilising the referenced task above. Background VRS 1063 is an existing AMM repair for the reattachment of LPC fan blade chocking pads, which can become detached during engine running and during removal of fan blades. Early standard blades have pads which are bonded to the fan blades using silicoset rubber compound and later standard blades have stick-on pads which use a double sided adhesive tape. The two configurations are called ‘Assembly A’ and ‘Assembly B’ in the repair. The purpose of the amendment to the repair is to include a procedure for a replacement of the stick-on pads and also to add the Uni-Directionally Profiled (UDP) blade part numbers 6A6519 for (A1) engines and 6A6521 for (A5/D5) engines.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Repair of Fan Blade Chocking Pads Note: There are two methods of attaching the chocking pads to the blade root: Assembly A AMM (72-31-11-300-019). Repair Scheme Number VRS 1063. This method uses primer for Silicoset (material no. V08-014) in conjunction with Cold Curing Silicone Compound (material no. V08-013). Essentially this method involves gluing the pads onto the fan blade. Assembly B AMM (72-31-11-300-019). Repair Scheme Number VRS 1063. This method uses double-sided adhesive tape to affix the pads to the blade.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Repair of the Fan Disk Rear Ramp During the removal operation of a fan blade, it is possible to dislodge the rear ramp from its location in the ‘dove-tail’ slot in the fan disk. Great care must be taken to inspect the fan disk and the security of the rear ramps, as they play an important role in providing a firm fixing and support for the individual fan blades. Should it be discovered that a rear ramp has become separated from the disk it must be refitted/replaced and a full description of the task can be found in the AMM task reference 72-31-12-300-010. This is summarised as follows: Remove the stage 1 fan blade from the stage 1 fan disk assembly Clean the disk and rear ramp bonding surfaces:

• Hand abrade the disk and rear ramp bonding area, using a scotch brite pad (material No. V05-126) or garnet paper (Material No. V05-017)

• Swab degrease the disk and rear ramp bonding areas, using a clean lint-free cloth made moist with methyl ethyl keytone (material No. V01-076)

Caution: Mating surfaces of the component must be scrupulously clean and contact surfaces must not be touched by hand or otherwise contaminated. Bonding must be carried out immediately following surface preparation

Bond the rear ramp to the disk:

• Apply masking tape to the rear ramp. Using masking tape (Material No. V02-019) Note! The masking tape is used in order to allow the engineer to hold and place the rear ramp accurately in the dovetail slot. See diagram on next page.

• Apply the adhesive to the disk and rear ramp bond areas. Use toughened acrylic adhesive with initiator (Material No. V08-114) Use a small spatula or trowel to apply the adhesive. Note The four ‘pips’ on the rear ramp, are to ensure adequate thickness of adhesive is maintained between the mating surfaces. See diagram on next page.

• Fix the rear ramp to the fan disk and remove the masking tape from the rear ramp.

• Use finger pressure to hold the rear ramp in position for three minutes.

• Cure the adhesive for one hour at room temperature between 21 deg. C. and 25 deg. C.

• Visually and dimensionally examine the bonded rear ramp.

• Install the stage 1 fan blade to the fan disk assembly.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance LP Compressor (Fan) Maintenance Repair of the Stage 1 Fan Disk After removal of a fan blade it is also necessary to carryout an inspection of the fan disk in accordance with the AMM task 72-31-12-200-010. This task includes the examination of the dovetail slots for peeling of the dry film lubricant. (Use a dental mirror in order to accomplish this. If there is any amount of peeling of dry film lubricant, carryout the repair VRS1149 in accordance with the AMM Ref task 72-31-12-300-011.This is summarised as follows: 1. Remove the stage 1 fan blade from the stage 1 fan disk

assembly: 2. Clean the missing coat areas on the dovetail slot:

• Use a lint-free cloth made moist with clean isopropyl alcohol (Material No. V01-124).

3. Apply the dry film lubricant to the dovetail slot of the disk:

• Touch up the dry film lubricant to the missing coat areas with a clean brush (Material No. V01-005), use multi-purpose high load dry lubricant (Material No.V10-005) or bonded lubricant (Material No. V10-106).

• Apply three coats of the dry lubricant to the total thickness of between 0.001 and 0.002 in (0.025 and 0.051mm) to surface ‘BJ’ (see diagram below).

4. Air dry as follows:

• If V10-005 multi-purpose high load dry lubricant was used, dry for 20 minutes

• If V10-106 bonded lubricant was used, dry for 30 minutes.

5. Visually examine the dry film lubricant on the dovetail

slot of the disk. Use a dental mirror:

• The layer must be smooth and bonded correctly to the surface of the part.

• The layer must not have any flakes or cracks. 6. Identify the repair:

• A log book entry is necessary when you have touched up the slot surface of over 50%. Write VRS1149 in the engine log book.





Fan Trim Balance Procedure

Additional information for reference use only Definitions: 1. Speed N1 expressed as a % - 100%= 5650 rpm 2. Amplitude ‘U’ indicated vibration level expressed in Mils

(P-P) ‘Peak to Peak’ from the Engine Vibration Monitoring Unit EMVU

3. Phase Angle ‘A’ indicated angle in degrees from the EMVU system.

4. Phase Lag ‘B’ dynamic phase lag of the LP system between phase angle and true position of unbalance.

5. Mass Coefficient ‘K’ value by which the phase amplitude must be multiplied to give correction mass for a given speed. Mils = American ‘Thousands’ of an inch Vibration is measured in “/sec i.e. velocity. Displacement is the movement of the casing when subjected to unbalanced loading effects: (a+b)= Total displacement (U) To carry out conversion use formulae Velocity = (U/2) * (rpm/60) * (2*pi) Worked example: given that U = 1.5 rpm = 5346 (approx. 95%)

Answer = (1.5/2)*(5346/60)*6.283 =419.9 mils/sec = 0.42 inches of movement per second

Out of balance point Rotation

(a) (b)

Revision 2 Page 4 -43

SECTION 5

ELECTRONIC ENGINE CONTROL

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Electronic Engine Control Electronic Engine Control Introduction The V2500 uses a Full Authority Digital Electronic Engine Control (FADEC). The FADEC comprises the sensors and data input, the electronic engine control unit (EEC) and the output devices, which include solenoids, fuel servo operated actuators and pneumatic servo operated devices. The FADEC 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 mounted unit, which is shielded and grounded as protection against EMI - 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 utilises 67.53 Watts of power from either the three phase AC from a dedicated engine mounted alternator, or 28 Volts DC from an aircraft source.

• Six ‘screened’ pressure ports provide the required pressure inputs to both channels.

• Built in handle facilitates removal and handling.

• Has three control modes in each channel. Engine Pressure Ratio (EPR) – which is the Primary thrust control Mode. N1 Rated and Un-rated and also provides Auto Starting and Thrust Reverser control. (To be covered in detail later).

• Schedules engine operation to provide maximum engine performance and fuel savings.

• Provides improved engine starting (Auto Start) and transient characteristics (acceleration/deceleration).

• Provides maximum engine protection and is more flexible to readily adapt to changes in engine requirements.

Revision 2 Page 5-1

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Electronic Engine Control

Revision 2 Page 5-2

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Electronic Engine Control The Engine Electronic Control (EEC) Description The EEC is a dual channel control unit that utilises a split housing design. The assembled unit is sealed with a housing seal and a protective shield provides channel separation. The control assembly is separated into two modules, each containing one control channel. Each module contains two multi-layer printed circuit boards assemblies, which enable it to function independently of the other channel. A mating connector provides ‘Crosstalk’, for partial or complete channel switching and fault isolation logic when the two modules are joined. This connector also provides for the exchange of ‘cross-link data’, cross wiring and hardwired discretes between the two channels. The EEC has two identical electronic circuits that are identified as Channel A and Channel B. Each channel is supplied with identical data from the aircraft and the engine. This data includes throttle position, aircraft digital data, air pressures, air temperatures, exhaust gas temperatures and rotor speeds. The EEC, to set the correct engine rating for the flight conditions uses this data. The EEC also transmits engine performance data to the aircraft. This data is used in cockpit display, thrust management and condition monitoring systems.

Each of the EEC channels can exercise full control of all engine functions. Control alternates between Channel A and Channel B for consecutive flights, the selection of the controlling channel being made automatically by the EEC itself. The channel not in control is nominated as the back up channel. .

Revision 2 Page 5-3


Revision 2 Page 5-4

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Electronic Engine Control Electronic Engine Control Harness (electrical) and Pressure Connections Two identical, but separate electrical harnesses provide the input/output circuits between the EEC and the relevant sensor/control actuator, and the aircraft interface. The harness connectors are 'keyed' to prevent misconnection.

Note: Single pressure signals are directed to pressure transducers - located within the EEC - 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 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

Electrical Connections

Front Face J1 E.B.U. 4000 KSA J2 Engine D202P 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

Revision 2 Page 5-5


Revision 2 Page 5-6

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Electronic Engine Control Engine Electronic Control (EEC.) Overview The EEC provides the following engine control functions:-

• Power Setting (E.P.R.).

• 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 logic built in. This logic operates continuously to detect and isolate defects in the EEC.

Revision 2 Page 5-7



© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Electronic Engine Control Electronic Engine Control EEC Interfaces The EEC interfaces with a number of other aircraft systems. The main systems are as follows:

Engine Interface Unit (EIU) Two EIU’s are fitted to the aircraft, the main functions are to:

• Supply aircraft data to the EEC. • Ensure engine to engine segregation. • Select aircraft electrical supplies to the EEC. • Supply data directly to other aircraft systems.

• Air Data Inertial Reference System (ADIRS) The main functions of the ADIRU’s are:

• To process pitot and static inputs. • Supply air data to other aircraft systems including EEC and

to the DMC’s for display.

Flight Warning Computer (FWC) Two FWC’s are fitted to the aircraft and their main function is to:

• Process data for fault annunciation. • Generate actions necessary for associated fault.

Display Management Computer (DMC) Three DMC’s are fitted to the aircraft and their main function are to:

• Receive and process data from other aircraft systems. • Format and display the data on the 6 display units. Flight Management and Guidance Computer (FMGC) Two FMGC’s are fitted and their main functions are:

• Flight Management, Navigation, performance optimisation and display management.

• Flight guidance, autopilot and thrust commands to the EEC.

Other aircraft systems interface with the EEC through the EIU. These are:

• Spoilers Elevator Computer (SEC).

• Landing Gear Control Interface Unit (LGCIU).

• Bleed Monitoring Computer (BMC)

• Flight Control Unit (FCU).

• Centralised Fault Display Interface Unit (CFDIU).

• Multipurpose Control and Display Unit (MCDU).

Revision 2 Page 5-9



© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Electronic Engine Control 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. 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 (Selected from multiple rating options).

• Engine Serial No.

Location The data entry plug is located on the channel B side electrical connectors of the EEC. During removal/replacement of the DEP it is necessary to use an EEC harness wrench, as it is imperative that the connectors are tight. On fitment of the DEP to the EEC align the main key of the

connector with the EEC and hand tighten the connectors. Then using the EEC Harness Wrench torque tighten the DEP connector to 32 lbf in. The DEP links the coded data inputs through the EEC by the use of shorting jumper leads which are used to select the plug pins in a unique combination. During the life of an engine, it may be necessary to change the DEP configuration, either during incorporation of Service Bulletins or after engine overhaul, when the EPR Modifier code may need to be changed. This is accomplished by changing the configuration of the jumper leads in accordance with the relevant instructions. Service Bulletin V2500-ENG-72-0285 contains the specific detail of the process involved for modifying the Data Entry Plug.









DATA ENTRY PLUG (DEP)

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Electronic Engine Control 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 pressures, (single pressure transducers within the EEC provide signals to each channel - A and B).

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

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

• An output fault in one channel will cause switchover to control from the other channel.

• In the event of faults in both channels a pre-determined hierarchy decides which channel is more capable of control and utilises that channel.

• In the event of loss of either channels, or loss of electrical power, the systems are designed to go to the fail safe positions.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Electronic Engine Control Failures and Redundancy In the event of loss of both input signals, loss of either channels, or loss of electrical power, the system is designed to go to the fail safe positions shown in the table below. EEC. System Component • Fuel Metering Unit

− Metering Valve Torque Motor

− Fuel Shut-off Valve

− Overspeed Valve Solenoid

• Seventh Stage Bleed Valves

• Tenth Stage Bleed Valve

• Combined Active Clearance Control Unit

− High ACC

− Low ACC

• Low Compressor (2.5) Bleed Actuator

• Stator Vane Actuator

• Note; If there is a failure of the thrust reverser control unit arming valve while the reverser is deployed, the. reverser will remain deployed.

Failsafe Position

• Minimum Fuel Flow Position

• Last Commanded Position

• Normal Fuel Flow Position

• Valves Open

• Valve Open

• Valve Closed

• Valve partially (-44%) Open

• Valve Open

• Vanes Open


© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Electronic Engine Control EEC. System Component

• Fuel Diverter and Back to Tank Valve.

− Fuel Diverter Valve.

− Fuel Back to Tank Valve.

• Air/oil Cooler Control Valve Actuator

• Tenth Stage “Make-up” Cooling Air Valve

• Thrust Reverser Control Unit.

• PT2/TT2 Relay Box

− ignition Relays

− Probe Heater Relays

• Starter Air Valve

• Anti-ice Air Valve

Failsafe Position

• Solenoid De-energised (Mode 4 or 5).

• Valve Closed - No Return to Tank (Mode 3 or 5).

• Valve Open.

• Valve Open.

• Reverser Stowed.

• Ignition ON.

• Heater OFF.

• Valve Closed.

• Valve Open


© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Electronic Engine Control Operation and Control EEC Power Supplies The electrical supplies for the EEC are normally provided by a dedicated alternator, which is mounted on and driven by the external gearbox.

Dedicated Alternator The permanent magnet alternator has two independent sets of stator windings that supply two independent, 3 phase frequency AC outputs to the EEC. These unregulated AC supplies are rectified to 28 volts DC within the EEC The Dedicated Alternator also supplies N2 signals for the EEC. This is provided by the frequency of a single phase winding in the stator housing as the ‘primary’ speed signal used by both Channels of the EEC and for the Flight Deck instrument display of engine actual speed. Should this signal fail, there is a ‘Back-up’ signal which is derived from one of the three phase windings of Channel ‘B’ power generation. There is no speed signal generation provided by the output of the coil windings of the Dedicated Alternator Channel ‘A’ power supply. 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.

• 28V DC is required for some specific functions, which include the thrust reverser, fuel on/off and ground

test power for EEC maintenance. During engine starts 28V DC is supplied from the aircraft bus bars until the dedicated alternator comes 'on line' at approximately 10% N2. Switching between the aircraft 28V supply and dedicated alternator power supplies is done automatically by the EEC so in the event of a total failure of the dedicated alternator the EEC is supplied from the aircraft 28V DC bus bars,




EEC POWER GENERATION

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Electronic Engine Control Dedicated Alternator Cooling Shroud Location It is important that the cooling shroud is orientated correctly for the differing variant engines. The shroud must be clamped with the arrow on the shroud aligned with the number ‘1’ indicated position for A5 and A1 applications. For D5 applications only the arrow on the clamp must align with the number ‘2’ indicated position. With the arrow aligned make sure that the dowel on the shroud engages in the adjacent cooling hole on the casing, this correctly aligns the cooling air inlet on the shroud with the cooling hole in the casing. Tighten the nut on the shroud to hold the shroud firmly in the correct position. Torque the nut on the alternator shroud to 180 – 220 lbfin (20 – 25 Nm) Connect the tube to the alternator shroud and torque tighten the tube nut to 283 – 310 lbfin (32 – 35 Nm) Safety the tube with locking wire. Connect the electrical connectors and torque them to 16 lbfin (1,8 Nm) If this is not carried out, the cooling airflow may not be able to enter the stator housing due to the cooling air hole on the stator being masked by the clamp body of the cooling shroud. The diagram below shows both arrangements and their relevant application.






DEDICATED ALTERNATOR (A5 CONFIGURATION)

SECTION 6

POWER MANAGEMENT

(Chapter 76)

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Power Management Power Management Purpose The power management system is designed to allow the control of engine power by either manual or auto throttle control. Location The aircraft throttle is located in the flight deck. This is in reference to the TLA resolvers. The EEC is engine intermediate case mounted. This is in reference 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 of the 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 as the Throttle Resolver Angle (TRA). The relationship between the throttle lever angle and the throttle resolver angle is linear therefore:

1 deg TLA = 1.9 deg TRA

Revision 2 Page 6 - 1

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Power Management


© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Power Management Throttle Control Lever Mechanism The throttle control mechanism for one engine is shown below. The control system consists of:

• The throttle control lever.

• The mechanical box.

• The throttle control unit. The throttle control lever movement is transmitted through a rod to the mechanical box. The mechanical box incorporates 'soft' detents, which provides selected engine ratings, it also provides "artificial feel" for the throttle control system. A second rod to the throttle control unit transmits the output from the mechanical box. The throttle control unit incorporates two resolvers and six potentiometers. Each resolver is dedicated to one EEC channel; the output from the potentiometers provides T.L.A. signals to the aircraft flight management computers. A rig pin position is provided on the throttle control unit for rigging the resolvers and potentiometers.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Power Management Bump Rating Push Button (A1 Engined Aircraft only) In some cases (optional) the throttle control levers are provided 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 EPR mode when aircraft is on ground. Bump Ratings can be de-selected at any time by actuating the bump rating push button, as long as the aircraft is on the ground and the Thrust Lever is not in the Max Take-Off detent. In flight, the bump ratings are fully removed when the Thrust Lever is moved from the Take-Off detent to or below 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 the associated EPR display. Use of Bump must be recorded. When one Bump button is selected, both engines will be at the "Bump Rated" value. Pressing Bump again deselects Bump Rating.

Flexible Takeoff (A1 & A5 Engine Aircraft) Definition of Flexible Takeoff: In many instances, the aircraft takes off with a weight lower than the maximum permissible takeoff weight. When this happens, it can meet the required performance with a decreased thrust that is adapted to the weight: This is called ‘Flexible Takeoff’ and the thrust is called ‘Flexible Takeoff Thrust’. The use of Flexible Takeoff Thrust saves engine life. The maximum permissible takeoff weight decreases as temperature increases, so it is possible to assume a temperature at which the actual takeoff weight would be the limiting one. This temperature is called ‘Flexible Temperature’ or ‘Assumed Temperature’ and is entered into the FADEC via the MCDU PERF TO page in order to get the adapted thrust. Note! If the thrust ‘Bump’ is armed for takeoff and flexible thrust is used, the pilot must use the Takeoff Performance determined for the non-increased takeoff thrust (without Bump).

• Thrust must not be reduced by more than 25% of the full rated thrust.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Power Management Throttle Control Lever Mechanism The throttle control lever moves over a range of 65 degrees, from minus 20 degrees to plus 45 degrees. An intermediate retractable mechanical stop is provided at 0 degrees. Forward Thrust Range The forward thrust range is from 0 degrees to plus 45 degrees.

• 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 off power (FLTO) at 35 degrees.

Reverse Thrust Range Lifting the reverse latching lever allows the throttle to operate in the range 0 degrees to minus 20 degrees. A detent at minus 6 degrees corresponds to thrust reverse deploy commanded and reverse idle power, minus 20 degrees is max reverse power. Auto Thrust System (ATS) The Auto Thrust System can only be engaged between 0 degrees and plus 35 degrees. Two engine operation 0 degrees and Max Climb 25 degrees. One engine operation 0 degrees and Max Cont 35 degrees.

Thrust Rating Limit Thrust rating limit is computed according to the thrust lever position. If the thrust lever is set in a detent the FADEC will select the rating limit corresponding to this detent. If the thrust lever is set between two detents the FADEC will select the rating limit corresponding to the higher mode.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Power Management EEC/Fuel System Interface Purpose To allow the throttle signal from the flight deck to be received by the EEC. The EEC will convert this signal into a fuel flow error in order to change the fuel flow for a power level change. Description Movement of the pilots throttle control lever is sensed by the dual resolvers, which signal the TRA to the EEC. The EEC computes the fuel flow, which will produce the required thrust. The computed fuel flow request is converted to an electrical current (I) which drives the torque motor in the Fuel Metering Unit (FMU) which modulates fuel servo pressure to move the Fuel Metering Valve (FMV) and sets the fuel flow. A dual resolver senses movement of the FMV, which is located in the fuel metering unit next to the FMV. The dual resolver translates the fuel metering valve movement into an electrical feedback signal that is fed back to the EEC.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Power Management Basic Control Loop The EEC uses closed loop control based on Engine Pressure Ratio (EPR) or, N1 if EPR is unobtainable. EPR Closed Loop Control The EEC computes a Target EPR as a function of:

• TRA (Throttle Resolver Angle).

• T amb (Ambient temperature).

• T2 (Engine air inlet temperature).

• Alt (Altitude).

• Mn (Mach Number). The EPR target is compared to the actual EPR to determine the EPR error. The EPR error is converted to a rate controlled fuel flow command (WF), which is summed with the measured fuel flow (WF actual) to produce the WF error. The W.F. error is converted to a current (I), which is sent to the FMU to drive the torque motor; this moves the FMV to change the fuel flow. The change in fuel flow causes the engine to accelerate/decelerate and brings about a change in actual EPR. This process continues until there is no EPR error. Note: The EEC controls the rate of change of fuel flow, and thus acceleration/deceleration times, as a function of the rate of change of HP Compressor Speed (N2).

N1 Reversion In case of no EPR (either sensed or computed) available, an automatic reversion to N1 mode is provided.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Power Management Alternate Engine Control N1 Reversion At the reversion to N1 mode (rated or unrated) an equivalent thrust to that achieved in EPR mode is provided until a thrust lever position change. Note: Autothrust control is lost. Rated N1 Mode An automatic reversion to rated N1 mode occurs if sensed EPR. (Either P2 or P4.9) or any of the computed EPR parameters are not available. A manual selection of N1 mode on the push buttons on the overhead panel selects rated N1 mode. (Note! The pilot is instructed to select N1 rated thrust on both engines). The N1 mode indication is displayed in blue on the upper ECAM. Also displayed, is the N1 rating limit corresponding to the thrust lever position. Note: This is not displayed in unrated N1 mode.

.





Alternate Engine Control N1 Reversion At the reversion to N1 mode (rated or unrated) an equivalent thrust to that achieved in EPR mode is provided until a thrust lever position change. With both engines now selected to N1 rated thrust control, they will now be controlled to provide the required thrust levels dictated by thrust lever physical position. As N1 mode is selected on both engines, the EPR indication of both engines is not available on the upper ECAM screen.

Unrated N1 Mode If in addition to losing EPR parameters either T2 or Altitude data is lost, then the EEC automatically reverts to unrated N1 thrust setting. The unrated N1 thrust setting requires the thrust to be set manually to an N1 speed. An over boost situation can occur in this mode at the full forward thrust lever position. The N1 mode indication previously displayed in blue on the upper ECAM is no longer displayed in unrated N1 mode. Note: This mode is a ‘non-dispatchable mode’.





Thrust Modes The engine operates in one of three thrust modes:

• Auto.

• Memo.

• Manual. Entering, exiting these three modes is controlled by inputs to the Engine Interface Unit (EIU) Auto Thrust Mode The auto thrust mode is only available between idle and MCT when the aircraft is in flight. After take off the throttle is pulled back to the max climb position, the auto-thrust system will be active and the Automatic Flight system will provide an EPR target to provide either: -

• Max climb thrust.

• An optimum thrust.

• A minimum thrust.

• An aircraft speed (Mach number) in association with the autopilot.

Memo Mode The memo mode is entered automatically, from Auto mode if;

• The EPR target is invalid.

• One of the instinctive disconnect buttons on the throttle is activated.

• Auto thrust is disconnected by the EIU. In the ‘Memo’ mode the thrust is 'frozen', to the last actual EPR value and will remain frozen until the throttle lever is moved manually, or, auto thrust is reset. Manual Thrust Mode This mode is entered anytime the conditions for AUTO or MEMO are not present. In this mode thrust is a function of throttle lever position.





Alpha Floor Protection If an aircraft stall is imminent, the ‘Auto Thrust System’ sets the engine power to maximum, regardless of actual throttle position. The thrust level that ‘Alpha Floor Protection’ provides is that equivalent to maximum EPR level at TOGA. Note: The conditions requiring ‘Alpha Floor Protection’ to be invoked are extremely rare, due to the circumstances which require its operation





Thrust Modes Manual Mode The engines are in the manual mode provided the A/THR function is:

• Not armed or

• Armed and not active (Thrust lever not in the operating range and no alpha floor).

In these conditions each engine is controlled by position of its thrust lever. The pilot controls thrust by moving the thrust lever between IDLE and TOGA positions. Each position of the thrust lever within these limits corresponds to an EPR. When the thrust lever is in a detent, the corresponding EPR is equal to the EPR rating limit computed for that engine. When the thrust lever is in the FLX/MCT detent On the ground

− The engine runs at the Flex takeoff thrust rating if the crew has selected a flex takeoff temperature on the MCDU that is higher than the current Total Air Temperature (TAT). Otherwise the engine produces Maximum Continuous Thrust (MCT)

Note! A change in FLEX TEMP during the takeoff has no effect on the thrust.

After Takeoff

− The pilot can change from the FLX to MCT by moving the thrust lever to the Take Off Go Around (TOGA) or Climb (CL) detent, then back to MCT. After that, they cannot use the FLX rating.

Note! Setting the thrust lever out of the FLX/MCT detent without reaching the TOGA or CL detent has no effect. The pilot can always demand Maximum Take Off thrust by pushing the thrust lever all the way forward, to the TOGA position.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Power Management Thrust Modes Automatic Mode In the Autothrust mode (A/THR function active), the FMGC computes the thrust, which is limited to the value corresponding to the thrust lever position (unless alpha-floor mode is activated




SECTION 7

FUEL SYSTEM (Chapter 73)

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System

Page 7-1 Revision 2

The fuel system controls are on the centre control pedestal and the indications are in the form of an annunciator light and ECAM messages.

Fuel System Introduction Purpose The primary purpose of the fuel system is to provide a completely controlled continuous fuel supply in a form suitable for combustion, to the combustion system. Description Control of the fuel supply is by the EEC via the Fuel Metering Unit (FMU). High pressure fuel is also used to provide servo pressure (actuator muscle) for the following actuators;

• BSBV actuators.

• VSV actuator.

• ACC actuator.

• ACOC actuator. The major components of the fuel system include;

• High and low pressure fuel pumps (dual unit).

• Fuel/oil heat exchanger.

• Fuel filter.

• Fuel metering unit (FMU).

• Fuel distribution valve.

• Fuel injectors (20).

• Fuel diverter and back to tank valve (FDRV).

Operation The aircraft pumps deliver the fuel to the engine LP pump. The LP pump boosts the initial fuel delivery to a pressure so as to prevent low pressure entry into the HP pump. Nominal pressure 150psi. The fuel flows into the fuel oil heat exchangers for the engine and IDG. Depending on the mode of operation the heat management system is in depends on which direction the fuel will flow. From the engine FCOC the fuel passes through the LP fuel filter. The filter has a 40 micron filtration capability. The fuel is received by the HP pump and is boosted to a nominal 1000 psi. The HP pump has pressure relief set at 1360 psi. The FMU meters the fuel and the excessive HP fuel is diverted back into the LP supply. The FMU is controlled by signals from the EEC. The fuel flow meter gives indication to the upper ECAM screen of real time fuel flow in KG/H. The distribution valve filters the fuel and splits the supply into ten separate outlets. The ten outlets supply fuel to two fuel spray nozzles per outlet. The fuel spray nozzles have small filters within them. This gives last chance filtration prior to fuel atomisation.


Revision 2 Page 7-2

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Controls The fuel metering valve position is controlled via the EEC from the movement of the thrust lever located on the centre control pedestal. The EEC has biased control of the FMU PRSOV for fuel selection to on and fuel selection to off, if N2 is below 50% and the start sequence is in auto. The command for fuel selection to off when the indicated N2 speed is above 50% is from the master lever. Indications The fuel temperature sensor is used by the EEC for the function of the heat management system. The fuel filter differential pressure switch annunciates to the lower ECAM screen a message of FILTER CLOG. This message is located in the right hand upper memo box. The message of FILTER CLOG will occur when the fuel filter differential pressure exceeds 5 psi. If there is a disagreement between the selection of the master lever and the PRSOV position then a fault exists.

Fuel valve failed to open: Master lever switch set to ON Fault light illuminates and master caution light illuminates accompanied by an audible tone. Upper ECAM message of;

ENG 1 FUEL VALVE FAULT -FUEL VALVE CLOSED Fuel valve failed to close Master lever switch set to OFF Fault light illuminates and master caution light illuminates accompanied by an audible tone. Upper ECAM message of;

ENG 1 FUEL VALVE FAULT -FUEL VALVE OPEN The fuel valve fault indications are inhibited during flight phase 3, 4, 7 and 8. Note: Amber caution with audible tone is a Class 1- Level 2 ECAM alert.

Revision 2 Page 7-3


Revision 2 Page 7-4

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel Pumps Purpose The fuel pumps are designed to ensure that the fuel system receives fuel at a determined pressure in order to allow the atomisation of fuel in the combustion chamber. Description The combined fuel pump unit consists of low pressure and high pressure stages that are driven from a common gearbox, output shaft. LP fuel pump Purpose To provide the necessary pressure increase to;

• Account for pressure losses through the Fuel Cooled Oil Cooler and the LP fuel filter.

• Suppress cavitation. • Maintain adequate pressure at the inlet to the HP

stage. Description Shrouded, radial flow, centrifugal impeller, with an axial inducer.

HP Stage Purpose To increase the fuel pressure to that which will ensure adequate fuel flow and good atomisation at all engine operating conditions. Description Two gear (spur gear) pump.

• Provides mounting for fuel metering unit (FMU). Integral relief valve.

Revision 2 Page 7-5


Revision 2 Page 7-6

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel Cooled Oil Cooler Purpose To transfer heat from the oil system to the fuel system to;

• Reduce the temperature of the engine lubricating oil under normal conditions.

Prevent fuel icing. Location The fuel and oil heat exchanger is located on the left hand side of the intermediate case. In the nine o’clock position. Description The fuel and oil heat exchanger is a single pass for the flow of fuel and multi pass for the flow of oil. The fuel and oil heat exchanger has the following features;

• A single casing houses the Fuel Cooled Oil Cooler and the LP fuel filter.

• Provides location for the fuel diverter and back to tank valve (unit not shown).

• Fuel temperature thermocouple.

• Fuel differential pressure switch.

• Oil system bypass valve. Fuel/oil tell tale leak indicator.

Revision 2 Page 7-7


Revision 2 Page 7-8


•

•

•

•

•

Low pressure Fuel Filter Purpose To remove solid contaminants from the LP part of the fuel system. Location The LP fuel filter is located in the LP fuel filter housing that is integral with the fuel and oil heat exchanger. Description The LP fuel filter is a woven, glass fibre, disposable, 40 micron (nominal) type. The LP fuel filter and housing have the following features;

• A differential pressure switch, which generates a flight deck message, FUEL FILTER CLOG, if the differential pressure across the filter, reaches 5 psi.

• A by-pass valve which opens and allows fuel to by-pass the filter if the differential pressure reaches 15 psi.

• A fuel drain plug, used to drain filter case or to obtain fuel samples.

• Fuel temperature sensor.

Service Bulletin ENG-79-0085 This Service Bulletin covers the fitment to engines of a FCOC incorporating design changes to prevent fuel leakage. A revised FCOC is introduced similar to the existing unit except for the following changes:

A revised fuel filter cap is introduced similar to the existing item except for increased doming and a change of material.

An adaptor is introduced between the FCOC housing and the fuel filter cap.

The fuel filter cap is attached to the new adaptor using 12 ‘D’ head bolts, locknuts and washers.

The adaptor is attached to the FCOC housing by five capscrews.

A small angular adjustment has been made to the circumferential position of the FCOC drain.

Revision 2 Page 7-9

© IAE International Aero Engines AG 2000 Fuel System


IAE V2500 Line and Base Maintenance

Pre-SB ENG-79-0085

SB ENG 79-0085 Fuel Filter Cap Adapter Views

LP FUEL FILTER CAP SB V2500-ENG-79-0085

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Purpose The FMU has three functions for fuel control. They are;

• Fuel metering to the combustion chamber.

• Control of the opening and closing off of the fuel supply to the combustion chamber.

• Overspeed protection. Location The FMU is mounted on the combined fuel pumps assembly. The combined fuel pumps assembly is located on the rear face of the high-speed gearbox, left hand side. Description The FMU is the interface between the EEC and the fuel system. All the fuel delivered by the HP fuel pumps, which is more than the engine requires is passed to the FMU. The FMU, under the control of the EEC, meters the fuel supply to the fuel spray nozzles. The HP fuel pressure also provides a servo operation (muscle) for the following actuators;

• Booster stage bleed valve (BSBV) actuators.

• Variable stator vane (VSV) actuator.

• Active clearance control (ACC) actuator.

• Air cooled oil cooler (ACOC) actuator.

Excessive HP fuel supplies that are not required, other than that for actuator control and metered fuel to the combustor, is returned to the LP system via the spill valve. In addition to the fuel metering function the FMU also houses the overspeed valve and the pressure raising and shut off valve. The overspeed valve under the control of the EEC provides overspeed protection for the LP (N1) and HP (N2) rotors. The pressure raising and shut off valve provides a means of isolating the fuel supplies to start and stop the engine. Note: There are no mechanical inputs to, or outputs from, the




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel Metering Unit (FMU) Service Bulletin V2500-ENG-73-0172 This Service Bulletin introduces a Woodward Governor Company FMU similar to the existing unit except for a ‘Common Flow/High Flow’ maximum fuel flow stop assembly. This allows the unit to be switched to suit all V2500-A5 model applications. This is considered logistically advantageous for mixed fleet operators. The changes introduced are:

a) The external single set fuel flow stop mechanism has been deleted.

b) An external switchable two-position maximum fuel flow stop has been introduced which can be set for either A319/A320 or A321 aircraft applications

c) A single reversible nameplate is introduced which, in conjunction with stop setting letter and FMU dataplate directive, will facilitate clear unambiguous identification of each flow setting.

d) A security seal system is introduced onto the above switchable fuel flow stop and reversible nameplate.

e) To facilitate installation of the security seal lock wire, the two existing retaining cap screws have been replaced by lockwire compatible equivalents.

FMU Part Number Position Setting Letter FMU 8061-636 0 FMU 8061-637 X

(i)To switch 8061-636 to 8061-637, carryout switch procedure in accordance with Woodward Governor Company Service Bulletin 83724-73 Fuel Metering Unit (FMU)

Service Bulletin V2500-ENG-73-0172 (Continued)

(ii) To switch 8061-637 to 8061-636, carryout switch procedure in accordance with Woodward Governor Company Service Bulletin 83724-73-0004.

a) Re-connect engine harness and LP fuel tube (Refer to AMM 73-22-52)

b) Close access to the engine (Refer to AMM 71-13-00)

c) Do an ‘idle’ check (Refer to AMM 71—00-00) or a wet motor leak test (Refer to AMM 71-00-00)

d) Do the operational tests of the starter and FMU (Refer to AMM 80-13-51)

Do the operational FADEC test as per (AMM 73-22-00)




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel Distributor Valve Purpose The fuel distributor valve receives fuel from the FMU and carries out three functions;

• Last chance filtration of the metered fuel.

• Distribution of the metered fuel through ten fuel supply tubes to the fuel spray nozzles.

• Upon shut down allows fuel drain back (pressure reduction) for prevention of fuel leaks into the combustor upon engine shut down.

Location The fuel distribution manifold is located on the right hand side of the combustion diffuser casing. It is in the 4 o’clock position. Description The fuel distributor manifold has the following features;

• Integral fuel filter - with by-pass valve.

• Single fuel metering (check) valve.

• Spring loaded closed upon engine shut down.

• Fuel pressure opened.

• Ten fuel outlet ports.

Fuel Distribution Manifold Purpose To allow the distribution of metered fuel from the distributor valve to the twenty fuel spray nozzles. Location The distribution manifolds are centred about the distributor manifold, they then branch out around the circumference of the combustion diffuser casing. Description There are ten distribution manifolds. Each manifold supplies fuel to two fuel spray nozzles. Note: The distribution manifold connectors have transfer tubes that allow a more positive seal to be achieved. If a leak is evident then it is prudent to suspect a seal failure.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel Spray Nozzles (FSN) Purpose To inject the fuel into the combustion chamber in a form suitable for combustion by;

• Atomising the fuel.

• Mixing it with HPC delivery air.

• Controlling the spray pattern. Location The fuel spray nozzles are equally spaced around the circumference of the combustor diffuser casing. Description Parker Hannifin manufactures the Airspray fuel nozzles. The fuel spray nozzles have the following features;

• 20 fuel spray nozzles.

• Inlet fitting houses fuel filter.

• Internal and external heat shields to reduce coking. Transfer tubes for improved fuel leak prevention.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel System Operation Fuel Metering Unit Description A simplified schematic representation of the Fuel Metering Unit is shown below. 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 (2) rotors.

• Isolation of fuel supplied for starting/stopping the engine.

Three valves arranged as follows carry out these three functions;

• The Fuel Metering Valve.

• The Overspeed Valve.

• The Pressure Raising and Shut Off Valve (PRSOV). Fuel metering valve The fuel metering valve varies the fuel flow according to the EEC command. The positional feedback to the EEC is by a rotary variable displacement transducer (RVDT). The overspeed valve The overspeed valve protects the engine against an exceedance of;

• N1 shaft speed.

• N2 shaft speed.

The feedback to the EEC of the valve operation is by a micro switch. The pressure raising and shut of valve (PRSOV) The PRSOV is an open and close type valve. The PRSOV controls the fuel to the combustor. When the valve is in the pressure raising state it is said to be open. When the valve is in the shut off state it is said to be closed. Note: The EEC has command to open the PRSOV upon an engine start. The EEC has command to close the PRSOV in auto start mode and when the N2 is below 50%. Above 50% N2 the close command is from the master lever in the flight deck only. Pressure drop governor and spill valve The pressure drop governor controls the pressure difference across the FMV. The spill valve is controlled by the pressure drop governor. The spill valve is designed to vary the excessive HP fuel pressure return to the LP system.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel System Operation Fuel metering Valve Operation Fuel metering is achieved by the Fuel Metering Valve and the Pressure Drop Regulator and Spill Valve, which act together in the following sequence: Signals from the EEC cause the torque motor to change position, which directs fuel servo pressure to re-position the Fuel Metering Valve. This changes the size of the metering orifice through which the fuel passes which in turn changes the pressure drop across the metering valve. The change in the pressure drop is sensed by the Pressure Drop Regulator which will re-position the spill valve and so increase/decrease the fuel flow through the fuel metering valve until the pressure drop is restored to its datum value. The increase/decrease in fuel flow causes the engine to accelerate/decelerate until the actual EPR is that demanded by the EEC signal. Movement of the Fuel Metering Valve is transmitted through a rack and pinion mechanism to drive a dual output position resolver. The resolver output is fed back to the EEC. The EEC automatically corrects changes in fuel density. Bi-metallic washers located in the pressure drop governor and spill valve assembly provide automatic compensation for changes in fuel temperature.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel System Operation FMU ‘Emergency’ Overspeed Protection The Overspeed Valve is positioned down stream, in series, with the Fuel Metering Valve. Note: It should be understood that this device is not incorporated to provide the usual N1/N2 red line limiting of max TO speed of 100%. The EEC will act through the Fuel Metering Valve to trim the fuel flow if N1 or N2 reach 100%. Operation The overspeed valve is spring loaded to the closed position, it is opened by increasing fuel pressure during engine start and during normal engine operation is always fully open. In the event of an overspeed condition (>109% N1 or >105.7% N2 ) the EEC sends a signal to the overspeed valve torque motor which changes position and directs HP fuel to the top of the overspeed valve thus 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 preventing the engine from being accelerated. The recommended procedure is for the flight crew to close the engine down, and not re-start. Closing down the engine is the only way to release the hydraulic latching.

Note: Because of the fact that the overspeed valve is spring loaded to the closed position, and opened by fuel pressure, the overspeed valve will close every time the engine is shut down. The PRSOV will remain open due to the small amount of fuel that is allowed to flow by the shutoff vale The microswitch gives valve positional feedback to the EEC.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System FMU ‘Emergency’ Overspeed Protection Should an ‘Emergency Overspeed’ condition be experienced and the engine has as a result exceeded 103% (either N1 or N2) the engine must be removed. Ref Task 71-00-00-991-156 figure 212 for N1 Ref Task 71-00-00-991-157 figure 213 for N2




© IAE International Aero Engines AG 2000 Fuel System IAE V2500 Line and Base Maintenance


Fuel System Operation Pressure Raising and Shut Off Valve Operation The third valve in the FMU is the Pressure Raising and Shut Off Valve (PRSOV). Its primary function is to isolate the fuel supplies to the fuel spray nozzles for starting and stopping the engine. It acts as a pressure raising valve to ensure that, during engine starts, fuel is not passed to the fuel spray nozzles until fuel pressures in the FMU are high enough to ensure the control devices will function correctly. The two position torque motor, which controls HP fuel pressure to operate the PRSOV also, controls a spill valve servo line. When the torque motor is selected to close the PRSOV, to shut down the engine, the spill valve servo line is opened. This will fully open the spill valve and direct all the HP fuel pump delivery back to the LP fuel system. The PRSOV torque motor is commanded open by the EEC during AUTO starts. It is commanded open by the engine master switch during a MANUAL start. The PRSOV can be commanded closed by the EEC during an AUTO start, if the EEC detects a fault in the start cycle.

The EEC’s ability to close the shut off valve is inhibited above 50% N2. Above 50% N2, and in flight, the PRSOV can only be closed by the crew operated switch in the flight deck. The microswitch gives valve positional feedback.



© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel System Maintenance The following line replaceable units are discussed for removal and installation. Note: The maintenance activities discussed are not intended for use during aircraft/engine activities. The AMM must be referenced in all cases. LP fuel filter Removal 73-12-42-000-010 Installation 73-12-42-400-010 LP/HP combined fuel pumps unit Removal 73-12-41-020-058 Installation 73-12-41-420-056 Fuel flow transmitter Removal 73-31-17-020-051 Installation 73-31-17-420-051 Fuel metering unit Removal 73-22-52-020-051 Installation 73-22-52-420-051 Fuel distribution valve Removal 73-13-43-020-010 Installation 73-13-43-400-010

Fuel distribution valve filter Removal 73-13-43-000-011 Installation 73-13-43-400-011 Fuel spray nozzle Removal 73-13-41-000-010

Installation 73-13-41-400-010





© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel Filter Element Removal/Installation The removal/installation details for the fuel filter is shown below. Reference must be made to the Aircraft Maintenance Manual. Removal Ch 73-12-42-000-010. Installation Ch 73-12-42-400-010. • Drain the fuel from the fuel cooled oil cooler by removing

the fuel drain plug into a suitable container.

• Remove the cap assembly from the FCOC. Note: Take care to remove the bolts gradually and smoothly in a symmetrical sequence so a single bolt is never subject to the total spring force on the cap. This also ensures that the components move apart gradually and do not become damaged. Removal Ch 73-2-42-000-010 • Remove and discard the fuel filter element.

• Inspect the cap assembly, FCOC mating faces, bolts and screws thread inserts.

• Discard all used gaskets and packing and replace with new items on reassembly.

Installation Ch 73-12-42-400-010 • Carefully install the new fuel filter element is in the correct

position in the FCOC housing.

• Install the cap assembly.

Note: Gradually tighten the bolts in asymmetrical sequence so a single bolt is never subjected to the total spring force on the cap.

• Lubricate the bolt threads with engine oil and torque load with limits quoted in the Aircraft Maintenance Manual.

• Do an idle leak check of the FCOC housing. Caution: Do not put fuel that has been drained from the engine back in to the fuel system. Warning: Be careful when you work on the engine components immediately after the engine is shutdown. The engine components can stay hot for up to one hour. Do not let engine fuel stay on your skin for a long time. Flush the fuel from your skin with water. The fuel is poisonous and can go through your skin and into your body. Do not let engine fuel or oil fall onto the engine. Unwanted fuel or oil must be removed immediately with a clean lint free cloth.

The fuel or oil can cause damage to the surface protection and to some parts.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System P/HP Fuel Pump Removal/Installation Reference must be made to the Aircraft Maintenance Manual. Removal Ch 73-12-41-020-058. • Remove the electrical connectors and the tubes from the

FMU (73-22-52).

• Remove the FMU from the fuel pump.

• Remove the fuel flow transmitter and associated tubes (73-31-17).

Note: It is necessary to remove the fuel flow transmitter to make sufficient space to remove the fuel pump. • Disengage the fail-safe latch of the clamp with a

screwdriver.

• Remove the stiff-nut from the T bolt.

• Remove the clamp.

• Remove the fuel pump. Caution: Hold the weight of the LP/HP fuel pump during removal. Clamp to prevent damage to the pump shaft and spline. (Weight 30.5 lb. (13 kg).


Installation Ch 73-12-41-420-056. • Install the pump on the adapter align the spline shaft of the

fuel pump with the spline gear of the gearbox. Caution: Hold the weight of the LP/HP fuel pump during installation. Clamp to prevent damage to the pump shaft and spline.

• Put the clamp around the flanges of the adapter and the fuel pump.

• Tighten the stiff-nut to engage with the fail-safe latch of the clamp.

Note: The fail-safe latch makes clicks when it is engaged.

• Torque load with limits quoted in the Aircraft Maintenance Manual.

• Install the fuel flow transmitter and associated tubes.

• Install the FMU to the fuel pump.

• Install the tubes and electrical connectors.

• Do idle leak check or wet motor leak check. Note:

Fuel pump failure can cause multi actuator failure.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel Flow Transmitter Removal/Installation Reference must be made to the Aircraft Maintenance Manual. Removal Ch 73-31-17-020-051 • Remove electrical connectors and tubes from the

transmitter.

• Remove the bonding lead.

• Remove the fuel flow transmitter.


Installation Ch 73-31-17-420-051 • Install the fuel flow transmitter.

• Install the bonding lead.

• Install electrical connectors and tubes to the transmitter.

• Torque load with limits quoted from the Aircraft Maintenance Manual.

• On upper ECAM DU check if fuel flow indication is available.

Do an idle leak check (71-00-00-710-046).




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel Metering Unit Removal/Installation Reference must be made to the Aircraft Maintenance Manual. Removal CH 73-22-52-020-051 • Remove electrical harness and raceway.

• Remove fuel tubes.

• Remove the FMU from the LP/HP fuel pump. Warning: Be careful during removal/installation of the FMU it weights 20 lb. (9 kg). • Discard all used gaskets and packing and replace with

new items on reassembly Caution: Some A319/A320 Aircraft require a specific part number FMU depending on the EEC installed. This is a certification requirement. Installation CH 73-22-52-52-420-051 • Install the FMU.

• Install the fuel tubes.

• Torque with limits quoted in the Aircraft Maintenance Manual.

• Install electrical harness and raceway.

• Do an operational test of the FADEC, no fault should show related to FMU.

Do an idle check (71-00-00710-012) or wet motor leak check

(71-00-00-710-046).




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel Distribution Valve Removal/Installation The removal/installation of the fuel distribution valve is shown below. Reference must be made to the Aircraft Maintenance Manual. Removal Ch 73-13-43-020-053 • Remove the fuel supply tube.

• Disconnect the fuel nozzle supply manifold nuts.

• Remove the fuel distribution valve.

• Remove the transfer tubes, the packing and the gaskets.


Installation Ch 73-13-43-420-010 • Install the transfer tubes and gaskets.

• Install the fuel distribution valve.

• Connect the fuel manifolds.

• Pressurise the system and check for leaks.

• Install the fuel supply tube.


• Do an idle check (71-00-00710-012) or wet motor leak check (71-00-00-710-046).

Fuel Distribution Valve Filter Removal/Installation Reference must be made to the Aircraft Maintenance Manual. Removal Ch 73-13-43-000-011 • Remove the fuel inlet line and the filter cover.

• Remove the filter. Installation Ch 73-13-43-400-011 • Install filter

• Install the fuel inlet line and filter cover.


Do an idle check (71-00-00710-012) or wet motor leak check (71-00-00-710-046).




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel Spray Nozzles Removal/Installation The removal/installation details for one of the fuel spray nozzles is shown below, the other 19 are similar but there are slight differences. Reference must be made to the Aircraft Maintenance Manual. Removal Ch 73-13-41-000-010 Installation Ch 73-13-41-400-010 The following points must be observed: -

• All the gaskets and packing used must be discarded on removal and replaced with new items on re-assembly.

• Lubricate the bolt threads with engine oil • Observe the torque loading limitations quoted in the

Aircraft Maintenance Manual. • Apply anti-galling compound (V10-032) to the shoulders of

the end fittings of the fuel supply tubes on re-assembly. • Apply white petrolatum (V10-041) to the rubber packing on

re-assembly. Note: Reference numbers e.g. V10-032 refer to consumable materials. A full list of these can be found in the Aircraft Maintenance Manual CH 70-30-00.

Warning Be careful when you work on the engine components immediately after the engine is shutdown. The engine components can stay hot for up to one hour. Do not let engine fuel stay on your skin for a long time. Flush the fuel from your skin with water. The fuel is poisonous and can go through your skin and into your body. Caution Do not let engine fuel or oil fall onto the engine. Unwanted fuel or oil must be removed immediately with a clean lint free cloth. The fuel or oil can cause damage to the surface protection and to some parts.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel System Control Harness The fuel system control harness electrical connections are shown below.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Fuel System Fuel System Harness The fuel system harness electrical connections are shown below.




SECTION 8

ENGINE OIL SYSTEM (Chapter 79)

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Engine Oil System Introduction Purpose The oil system is a self contained, full flow recirculating type design to ensure reliable lubrication and cooling under all circumstances. Description Oil cooling is controlled by a dedicated Heat Management System which ensure that engine oil, IDG oil and fuel temperatures are maintained at acceptable levels while ensuring the optimum cooling configuration for the best engine performance. The engine oil system can be divided into three sections. These sections are:

• Pressure feed.

• Scavenge.

• Venting. Pressure feed The pressure feed system uses the full flow generated by the pressure pump. The pressure pump moves the oil through:

• The pressure filter.

• Fuel oil heat exchanger. The oil is then distributed to the engine bearings and gear drives.

Scavenge The scavenge system is designed to retrieve the oil that is present in the bearing chambers and gearbox for cooling and recirculation. There are six scavenge pumps that are designed to suck the oil and pass it through:

• Magnetic chip detectors.

• A scavenge filter and master chip detector. Prior to returning the oil back to the oil tank. Venting The venting system is designed to allow the air and oil mix that develops in the bearing chambers and gearbox to escape to the de oiler. No.4 bearing does not have a scavenge pump. It relies upon the build up of air pressure in the bearing chamber to force the air and oil through the no.4 bearing scavenge valve and into the de oiler. Indications There are flight deck indications that allow the oil system to be monitored. There are also messages generated ECAM for further flight crew awareness.

Revision 2 Page 8-1

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System

Revision 2 Page 8-2

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Engine Oil System Indications The operation of the engine oil system may be monitored by the following flight deck indications: • Engine oil pressure. • Engine oil temperature. • Oil tank contents. In addition ECAM alerts may be given for the following non-normal conditions: - • Low oil pressure. • Scavenge filter clogged or partly clogged (high

differential pressure). • No 4 compartment scavenge valve inoperative. The oil system parameters are displayed on the Engine page on the Lower ECAM screen. Oil temperature (deg.c) Normal indication to ECAM is GREEN. 156°C or above flashing green indication. 156°C or above more than 15 minutes or 165°C without delay steady amber indication. Upper ECAM message ENG 1(2) OIL HI TEMP-Class 1, Level 2. Single chime. Master caution light Oil low temperature alert; throttle above idle and engine running. Upper ECAM message ENG 1(2) OIL LO TEMP-Class 1,

level 2. Single chime. Master caution light. Oil quantity Normal indication to ECAM is GREEN. Less than 5 quarts flashes green. Oil pressure Normal indication to ECAM is GREEN. 390 psid or above indication flashes. 60-80 psid amber indication. Upper ECAM amber message ENG OIL LO PR Class 1, level 1. 60 psid or below red indication. Master warning light. Continuous repetitive chime. Upper ECAM red message Class 1, level 3: ENG 1(2) OIL LO PR THROTTLE 1(2) IDLE Scavenge filter clog If the filter differential pressure is greater than 12 psi oil filter clog message appears on Engine page, lower ECAM. Oil Consumption Acceptable oil use is not more than 0.6 US pts/hr (0.5 Imp pts/hr). Oil increase of 100 ccs or more analyse sample for fuel contamination.

Revision 2 Page 8-3


Revision 2 Page 8-4

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Oil System Bearings and Gears Lubrication Front Bearing Compartment (Bearings no. 1, 2, 3) Purpose Bearings and gears require oil for:

• Lubrication.

• Cooling.

• Vibration suppression. Location The following bearings and gears are located in the front bearing compartment:

• Ball bearing no.1. (LP Thrust)

• Roller bearing no.2. (LP Radial)

• Ball bearing no.3. (HP Thrust) Description The bearing chamber utilises 1 hydraulic seal and 2 carbon seals to contain the oil within the bearing chamber. The front and rear seal of the LPC booster has stage 2.5 air passing across the seals in order to prevent oil loss. The hydraulic seal has HPC8 air passing across the seal in order to prevent oil loss between the two rotating shafts. The bearings and gears are fed with oil by utilising oil jets that liberally allow oil to enter the bearing area. The front bearing compartment has:

• Oil fed from the pressure pump.

• Scavenge oil recovery by the scavenge pumps.

• Vent air outlet to allow the sealing air to escape to the de oiler.

Revision 2 Page 8-5


Revision 2 Page 8-6

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Oil System Bearings and Gears Lubrication Centre Bearing Compartment (Bearing no.4) Purpose Bearings require oil for:

• Lubrication.

• Cooling. Location The following bearing is located in the centre bearing compartment:

• Roller bearing no.4. (HP Radial) Description The centre bearing compartment is the hottest compartment in the engine. In order to maintain the bearing at an acceptable operating temperature HPC12 air is taken from the engine, it is cooled by an air cooled air cooler (ACAC) and passed back into the engine. This cooling and sealing air is called buffer air. The buffer cooling air supply flows around the outside of the bearing in a cooling type jacket. In addition to cooling the buffer air is allowed to pass across the carbon seal and pressurise the no.4 bearing. This bearing compartment has the following:


• Scavenge oil and vent air recovery by the build up of

pressure in the bearing compartment forcing the air and oil out. The air and oil passes through the no.4 bearing scavenge valve and then into the de oiler.

Revision 2 Page 8-7


Revision 2 Page 8-8

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Oil System Bearings and Gears Lubrication Rear Bearing Compartment (Bearing no.5) Purpose Bearings require oil for:

• Lubrication.

• Cooling.

• Vibration suppression. Location The following bearing is located in the rear bearing compartment:

• Roller bearing no.5. (LP Radial) Description The rear bearing compartment has one carbon seal. This seal allows HPC8 air to leak across the seal thus preventing oil loss from the bearing compartment. This bearing compartment has the following:


• Scavenge oil recovery by the scavenge pumps. There is no vent outlet. The vent air is removed from the bearing compartment along with the scavenge oil.

Revision 2 Page 8-9




Page 8-11

on 2 • Oil jets directing the oil onto the gears.

Oil System Bearings and Gears Lubrication High speed external gearbox Purpose Gears require oil for:

• Lubrication.

• Cooling.

• Vibration suppression. Location The following module is located at the six o’clock position on the intermediate module. Description The high speed external gearbox is a one piece casting consisting of the following;

• Gear trains.

• Oil jets.

• Two scavenge outlets with strainers.

• Vent out to the de oiler.

• Integrally mounted oil tank.

• Angle gearbox.

• Mounting pads for the accessory units. The gear ratios differ to suit the rotational operating speeds of the accessory units. The high speed external gearbox gears are lubricated by:

• Splash lubrication caused by the motion of the gears. The high speed external gearbox has:


• Scavenge oil recovery by two scavenge pumps.

• Vent air outlet to allow the vent air to escape to the de oiler.

Revisi



© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Oil Tank Purpose To store the dedicated engine oil supply. Location Located to the top LH side of the external gearbox. Description The engine oil tank has the following features: Pressurised hot tank. Oil quantity transmitter.

• Gravity fill port with safety flap.

• Sight glass oil level indicator.

• Internal 'cyclone' type de aerator.

• Tank pressurisation valve (6 psi) ensures adequate pressure at inlet to oil pressure pump.

• Strainer in tank outlet to pressure pump.

• Provides mounting for scavenge filter and master magnetic chip detector (MCD).

The oil tank has the following for oil capacity:

• Tank capacity is 29 US quarts.

• Usable oil 24 US quarts. There is an anti siphon tube that supplies a small flow of oil back to the tank. This flow of oil splashes across the sight glass providing a cleaning action that prevents the build up of impurities.

On early A1 engines the oil tanks were fitted with a Prismalite oil level indicator, no sight glass was fitted.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Pressure Pump and Pressure Filter Assembly Purpose The pressure pump is designed to produce oil pressure for distribution in the bearing chambers and high speed external gearbox. The pressure filter gives initial filtration of the oil as it leaves the oil tank. Location The pressure pump and filter are one assembly. They are located on the front face of the high speed external gearbox. Mounted to the left hand side. Description The pressure pump and filter assembly has the following features:

• Cold start pressure limiting valve.

• Flow trimming valve (Line adjustment not permissible).

• Pressure Filter, 125 micron filtration.

• Combined filter bowl drain and priming port.

• Anti drain valve.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Air/Oil Heat Exchanger (Air Cooled Oil Cooler) Purpose The air cooled oil cooler acts as a second cooler for the oil system. The heat management system of the EEC controls the operation of this unit. Location Attached to the fan casing on the lower RH side. Description The air oil heat exchanger is normally closed when the engine fuel and oil temperatures are operating within their required temperature ranges. During certain conditions of engine operation the fuel and oil systems may experience high temperatures. In this condition, the air cooled oil cooler cools the oil in order for the oil to cool the fuel. The air cooled oil cooler has the following features:

• Corrugated fin and tube with a double pass design.

• Oil by pass valve.

• Modulated airflow as commanded by EEC (heat management system) Airflow regulated by air control valve.

• Electro-hydraulic servo valve operated system.

Note: The oil has a continuous flow through the air cooled oil cooler. This is regardless of whether the valve is open or closed. The Fuel Cooled Oil Cooler (FCOC), also referred to as the Fuel/Oil Heat Exchanger (FOHE), carries out oil system primary cooling.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Fuel/Oil Heat Exchanger Purpose The purpose of the fuel oil heat exchanger is to:

• Cool the engine oil and heat the fuel for most conditions, or -

• To use oil that has been cooled by the air cooled oil cooler (ACOC), to cool the fuel when it is too hot.

Location The location of the fuel oil heat exchanger is on the left hand side of the intermediate case. Description The fuel oil heat exchanger is a;

• Single pass fuel flow.

• Multi pass oil flow. Forms an integral unit with the LP fuel filter. A pressure relief valve permits oil to by-pass part of the cooler if the oil pressure is high during initial engine running, following a cold start.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Scavenge Pumps Unit Purpose Returns scavenge oil to the tank. Location All 6 scavenge pumps are housed together as a single unit on the rear of the high speed external gearbox, left hand side. Description The scavenge pumps assembly consists of six gear type pumps that are designed to retrieve the oil from the gearbox and bearing chambers. Thus returning the oil back to the oil tank. As all the pump gears are the same diameter, the scavenge pump capacity is determined by the gear width.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System De oiler Purpose To separate the air and oil mixture that develops in the bearing compartments and gearbox. To return the oil back to the oil tank and eject the air overboard. Location The de oiler is located on the front face of the high speed external gearbox, right hand side. Description The de oiler has the following features:

• Provides mounting for the No.4 bearing chamber scavenge valve.

• Overboard vent.

• Provides location for No.4 bearing Magnetic Chip Detector housing.

The de oiler is also called a centrifugal separator. This due to the fact that it relies upon the high rotational speed to centrifuge out the heavier weight oil from the lighter weight air. The oil is centrifuged outwards and into the gearbox and then scavenged back to the tank. The air is forced inwards by weight of continuos, flow through the rotor discharge slots, and then overboard via the vent pipe. .




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Scavenge Filter Purpose To trap solid contaminants. Location Mounted to the rear of the oil tank. Description The scavenge filter has the following features;

• The filter is a disposable mesh type filter with a nominal 30 micron filtration capability.

• A differential pressure switch monitors the flow through the filter for of blockage by contamination.

• A by pass valve, which opens when the filter flow is restricted due to contamination.

• The master magnetic chip detector housing is located on the filter housing.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System No 4 Bearing Scavenge Valve Purpose Maintains the centre bearing compartment (No 4 bearing) seal deferential pressure by controlling the venting of the compartment air/oil mixture to the de oiler. Location The no.4 bearing scavenge valve is located on the front of the de oiler, which is located on the front face of the external gearbox. Description The no.4 bearing scavenge valve has the following features;

• Operational feed back signal to EIU.

• Uses HPC10 air as the servo air for the valve operation.

• Stage 10 air less than 150 psi the valve is at maximum open position.

• Stage 10 air more than 200 psi the valve is at minimum open position.

• Feedback to EIU of valve operation is the valve position indicator; scavenge oil pressure sensor and Pb indication from the EEC.

The no.4 bearing scavenge valve controls the flow of the scavenge oil and vent air by varying the size of the orifice of the valve. This allows the scavenge oil and vent air to enter the de oiler under controlled conditions.

High flow When the engine is at low power the valve is at the high flow position. Therefore the valve is fully open and the pressure differential is maintained across the carbon seal. Low flow When the engine is at high power the valve is at the low flow position. Therefore the valve is at the restricted flow condition and the pressure differential is maintained across the carbon seal. Note: High flow at high power will cause a lower seal differential pressure. This will lead to the flow of buffer air across the carbon seal to increase. The increase flow of buffer air leads to the carbon seal drying out. If the valve is not in the correct position for the engine power setting, the warning message “BEARING 4 OIL SYS” on the upper ECAM.


IAE V2500 Line and Ba

Revision 2

se Mainte

LOW FLOW HIGH POWER

HIGH FLOW LOW POWER HPC1

A

NO.4 BEARING SCAVENGE VALVE

© IAE International Aero Engines AG 2000 Engine Oil System

Page 8-28

nance


COOLED HPC12 AIR IN (BUFFER AIR)


OUTLET TO DE OILER

0 SERVO IR IN

NO4 BEARING SCAVENGE OIL AND VENT AIR OUT TO THE N0.4

BEARING SCAVENGE VALVE.

DET

V250

373

VALVE POSITIONAL FEEDBACK TO EIU

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Magnetic Chip Detectors (MCD) Purpose The magnetic chip detectors (MCDs) allow on condition monitoring of the gears and bearing assemblies. Location The MCDs are located about the high speed external gearbox, as shown below. Description A total of 7 MCD’s are used in the oil scavenge system. The MCD’s have the following common features:

• The MCDs are of a bayonet style.

• Dual seal rings.

• Baulking pin preventing complete insertion in the case of a missing seal.

• Self sealing and removable housings. Each bearing compartment and gearbox has its own dedicated MCD (two in the case of the main gearbox). The No4 bearing is located in the de-oiler scavenge outlet. The master MCD is located in the combined scavenge return line, at the scavenge filter inlet. If the master MCD indicates a problem then each of the other MCD’s are inspected to indicate the source of the problem. Access to the dedicated MCD’s is by opening the L and R hand fan cowls.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Master Magnetic Chip Detector Purpose The master MCD gives indication for all gears and bearings of the engine. It allows periodic inspection without the requirement to inspect all MCDs. Location The master MCD can be accessed from a dedicated panel on the left hand side fan cowl door. Description The master magnetic chip detector is located in the scavenge filter case at the inlet to the filter. If debris is found on the master MCD, the source can be determined by inspecting all the MCDs. Further confirmation can be had by inspecting the scavenge filter.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Magnetic Chip Detectors Left Hand Side Location The following is the location for the;

• No’s 1, 2 and 3 bearings MCD.

• Main gearbox left hand side scavenge pick up MCD.

• Angle gearbox MCD. Are located to the rear of the main gearbox on the left hand side.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Magnetic Chip Detectors Right Hand Side Location The following is the location for the;

• No.5 bearing.

• De oiler (No.4 bearing).

• Main gearbox right hand scavenge pick up.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Differential Oil Pressure Transmitter and Low Oil Pressure Warning Switch Purpose The pressure transmitter is designed to give real time indication to the ECAM of differential oil pressure. The low pressure switch is designed to give warning of minimum operating differential oil pressure to ECAM. Location The differential oil pressure transmitter and low oil pressure switch are located on the left hand side intermediate case. Located in the 10 o’clock position. Description The pressure transmitter and low oil pressure switch differential pressures are sampled from:

• Pressure feed to the no.4 bearing.

• Scavenge oil from the no.4 bearing.


© IAE International Aero Engines AG 2000 Engine Oil System

Revisi

on 2 Page 8-38

PRESSURE TRANSMITTER AND LOW PRESSURE SWITCH

IAE V2500 Line and Base Maintenance

DETV252106

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Oil System Maintenance The following oil system components are discussed for maintenance. Check oil level AMM ref. 12-13-79-610-011 MCD Inspection AMM ref. 79-00-00-601 No.4 Bearing Scavenge Valve Removal AMM ref. 79-23-51-000-010 Installation AMM ref. 79-23-51-400-010 Pressure Filter Removal AMM ref. 79-21-44-000-010 Installation AMM ref. 79-21-44-400-010 Scavenge Filter Removal AMM ref. 79-22-44-000-010 Installation AMM ref. 79-22-44-400-010 Oil Scavenge Pump Removal AMM ref. 79-22-43-000-010 Installation AMM ref. 79-22-43-400-010 Full reference must be made to the AMM.

Check oil level AMM ref 12-13-79-610-011. Note: check oil level 30 minutes after engine shut down. If the engine has been shut down for more than 1 to 10 hours run the engine at idle for a least 3 minutes. If the engine has been shut down more than 10 hours the engine has to be dry cranked then run at idle for a least 3 minutes.

• Open oil servicing panel in the left fan cowl.

• Remove filler cap from the engine oil tank.

• Fill the engine oil tank – gravity fill or pressure fill.

• Make sure oil level sight glass shows “FULL”. Note: do not fill the oil tank past the sight glass “FULL” level. Filling to tank overflow will result in excess oil, leading to amber cross indication warnings and service disruption.

• Install the oil tank filler cap.

• Close the engine oil tank servicing access panel




Notch ‘3’ 20 litres 22 US Quarts 4.5Imp Gal

Notch ‘2’ 23 litres 24 US Quarts 5.1 Imp Gal

Notch ‘1’ 26 litres 27 US Quarts 5.7 Imp Gal

‘Full’ level notch 27.3 litres

29.0 US Quarts 6.0 Imp Gal

(Within 30 minutes from engine shutdown

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Oil System General Magnetic Chip Detectors Inspection/Check AMM ref. 79-00-00-601 It is recommended that records be kept of all debris found on the MCD’s and in the filters. The debris must be examined with a binocular microscope of 20 times magnification. If the engine or external gearbox is rejected, the debris from the MCD’s and filters must be sent with the component to the overhaul shop. There are four main types of debris found on MCD’s and in filters. They are build debris, magnetic fines, metallic flakes and chips and gear tooth fragments. Build Debris This is unwanted material that is accidentally left in the engine when it is assembled. The build debris comes from machining operations when the components are manufactured. When the material is broken it can look the same as gear or steel seal material. Magnetic Fines These are very small steel particles, which show as a black sludge on the MCD’s. When the oil is removed they show as dull hair like slivers. Fines come from permitted engine wear. Bearing skid can also make fines, but the increased quantity will show during the analysis of MCD samples.

Metallic Flakes Metallic flakes come from these components: ball bearings, roller bearings and gear teeth. Flakes that have an irregular shape must be examined to find their origin.

• Ball bearing and ball bearing track flakes are usually almost circular with radial cracks. When the flake is clean, the shiny side is much brighter than other types of flake. The shiny side also has small scratches that go across each other.

• Roller bearing and roller bearing track flakes are usually almost rectangular in shape. When the flake is clean, the shiny side is much brighter than other types of flake. The shiny side has small scratches that go across each other.

• Gear teeth flakes are shiny with an irregular shape. They are usually thicker and not as bright as ball or roller bearing flakes.

Chips These are very thick flakes or pieces of metal, which usually have one smooth machined surface. Gear Tooth Fragments These are the corner pieces of gear teeth and may show that the gears are not correctly aligned.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System No4 Bearing Scavenge Valve Removal/Installation Reference must be made to the Aircraft Maintenance Manual. Removal AMM ref. 79-23-51-000-010 • Carry out the necessary safety precautions in the flight

deck.

• Open fan cowl doors.

• Remove the electrical connector from the valve.

• Remove the pressure sensor tube.

• Remove the pneumatic supply tube.

• Remove the scavenge oil tube.

• Remove the no.4 bearing scavenge valve. Installation AMM ref. 79-23-51-400-010 The installation procedure is the reverse of the removal procedure. If the original valve is to be installed then all seal rings must be replaced.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Pressure Filter Removal/Installation Reference must be made to the Aircraft Maintenance Manual. Removal AMM ref.79-21-44-000-010. • Remove the nut from the cover.

• Drain the filter casing.

• Remove the filter cover.

• Remove the filter element.

• Discard all used gaskets and packing and replace with new items on re assembly.

Note: The pressure filter can be cleaned ultrasonically. Installation AMM ref. 79-21-44-400-010. • Install the filter element.

• Install the cover.

• Install the nut to the cover.

• Torque load bolts with limits quoted in the Engine Maintenance Manual.

• Fill the engine oil system as necessary (12-13-79-610-011).

• Do an idle leak check (71-00-00-710-012) or do a dry motor leak check (71-00-00-710-045).

Warning: Be careful when you work on the engine components immediately after the engine is shutdown. The engine components can stay hot for up to one hour. Caution: Do not put oil that has been drained from the engine back in to the oil system. Do not let engine fuel or oil fall onto the engine. Unwanted fuel or oil must be removed immediately with a clean lint free cloth. The fuel or oil can cause damage to the surface protection and to some parts.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Pressure Pump/Filter Housing Removal/Installation Reference must be made to the Aircraft Maintenance Manual. Removal AMM ref.79-21-41-000-010. • Drain the oil tank.

• Disconnect the air cooled oil cooler oil supply tube.

• Remove six nuts. Caution: Hold the weight of the oil pressure pump and filter assembly during the removal of the nuts.

• Remove the pump assembly


Caution: Make sure that the sleeve and suction strainer do not fall during the removal of the oil pressure pump Use a tool with a blunt edge (for example a smooth putty knife) to separate the pump from the gearbox flange. You must make sure that you do not damage the sealing surfaces of the pump piloting diameter. Installation AMM ref. 79-21-41-400-010. • Install new packings.

• Install sleeve and suction strainer.

• Install the oil pump and filter assembly.

• Connect the air cooled oil cooler supply tube.


• Fill the engine oil system as necessary (12-13-79-610-011).


•




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Scavenge Filter Removal/Installation The removal/installation details for the oil scavenge filter is shown below. Reference must be made to the Engine Maintenance Manual. Removal AMM ref. 79-22-44-000-010 • Drain the scavenge oil filter casing by the drain plug.

• Remove the scavenge oil filter cover.

• Remove the filter element from the filter housing. Note: The scavenge filter is not re-usable.


Installation AMM ref. 79-22-44-400-010 • Install two guide pins (IAE 1F10082) onto the filter

housing.

• Install scavenge filter element.

• Install filter cover.

• Attach filter cover with three bolts and washers by hand.

• Remove the two guide pins.

• Install the remaining bolts and washers.


• Install drain plug.


Service Tip Temporary “engine oil filter clog” messages may be triggered at oil temperatures below 10 °C (50 °F), generally during the first start of the day. No maintenance action is required if an “Oil Filter Clog” message displays temporary in the aircraft deck prior to engine achieving a stabilised idle condition.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Oil Scavenge Pump Removal/Installation Reference must be made to the Aircraft Maintenance Manual. Removal AMM ref. 79-22-43-000-010 • Drain the oil system.

• Remove scavenge oil filter housing.

• Remove scavenge oil tubes from the scavenge oil pump.

• Disconnect the adapter from the angle gearbox.

• Remove the scavenge oil pump. Warning: Be careful during the removal of the scavenge oil pump it weighs 12.22 lb. (5.544 kg). Caution: Hold the weight of the scavenge oil pump during removal of the nuts. Make sure that the suction strainer does not fall during the removal of the scavenge oil pump. Use a tool with a blunt edge (for example a smooth putty knife) to separate the pump from the gearbox flange. You must make sure that you do not damage the sealing surfaces of the pump piloting diameter.

• Remove the suction strainer.

Installation AMM ref. 79-22-43-400-010 • Discard all used gaskets and packing and replace with

new items on re assembly.

• Align the scavenge oil pump and the gearbox drive without the packing.

Caution: Do not turn the engine, gearbox gear train, scavenge pump drive gear. This is to avoid damage to the pump/gearbox during re-installation of the pump. Note: Turn the gearbox gear train with the wrench. This helps to engage the pump drive gear in the gearbox. The position of the wrench will show if the engine/gearbox gear train rotation occurred after removal of the pump.

• Remove the scavenge oil pump.

• Install the suction strainer.

• Install the scavenge oil tubes.

• Install the scavenge oil pump with packing.

• Connect the adapter to the angle gearbox.

• Install the scavenge oil tube to the scavenge oil pump.

• Install the scavenge oil filter housing.

• Install the crank cover.

• Fill the oil system. Do an idle leak check (71-00-00-710-012) or do a dry motor leak check (71-00-00-710-045).




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System Oil System General Inspection/Check Oil Consumption A survey was carried out on oil consumption taken over three weeks, surveying 58 engines results as follows: - Average oil consumption 0.12 US Pints/Hour. Range, engine to engine 0.04 to 0.14 US Pints/Hour. In service consumption Fleet range 0.11 to 0.14 US Pints/Hour. In service consumption A1 Fleet range 0.20 to 0.42 US Pints/Hour. Oil Quantity These figures should be used as guideline figures only rather than strict limits. Minimum oil quantity cockpit indications Before start. 11 quarts + estimated consumption during flight. (Max average consumption 0.3 qts/hr.).

Engine at ground idle. 6 quarts + estimated consumption during flight (Max average consumption 0.3 qts/h.) Oil Condition There are three types of oil condition to consider: - Note: The colour of new oil can be different between oil brands. Some oil brands are dark when delivered as new. In general, the oil colour will get darker with engine operation. Black oil Oil that has suspended particles of carbon and appears usually dark or black in colour. Contaminated oil Oil, which is contaminated with foreign substances such as, hydraulic fluid, fuel etc. Degraded oil Oil that has undergone physical property changes (viscosity, acidity, etc). Refer to Aircraft Maintenance Manual for Total Acid Number (TAN) tests and oil viscosity tests. Total Acid Number given in mg/g (KOH) and viscosity usually given in Centistokes (cSt) at 100 deg. C.



This page is left intentionally blank


© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Oil System The oil system harness electrical connectors are shown below.




SECTION 9

HEAT MANAGEMENT SYSTEM

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Heat Management System Heat Management System Purpose The system is designed to provide adequate cooling, to maintain the critical oil and fuel temperatures within specified limits, whilst minimising the requirement for the fan air offtake. Location The following units are located about the engine fan case;

• The engine FCOC.

• The engine ACOC.

• The IDG FCOC.

• The fuel diverter and back to tank valve.

• The aircraft outer wing fuel tanks. Description Three sources of cooling are available: -

• The LP fuel passing to the engine fuel system.

• The LP fuel that is returned to the aircraft fuel tanks.

• Fan air.

There are four basic configurations between which the flow paths of fuel in the engine LP fuel system are varied. The configurations are;

• Mode 1.

• Mode 3.

• Mode 4.

• Mode 5. Within each configuration the cooling capacity may be varied by control valves that form the fuel diverter and back to tank valve. The transfer between modes of operation is determined by software logic contained in the EEC. The logic is generated around the limiting temperatures of the fuel and oil within the system together with the signal from the aircraft that permits/inhibits fuel spill to aircraft tanks.

Revision 2 Page 9-1

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Heat Management System

Revision 2 Page 9-2

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Heat Management System Air/Oil Heat Exchanger Air Modulating Valve Purpose To govern the flow of cooling (fan) air through the air/oil heat exchanger, as commanded by the EEC heat management control system. Location The air/oil heat exchanger (ACOC) is attached to the right hand side of the engine fan case. It is in the four o’clock position as viewed from the rear of the engine looking forwards. Description The ACOC is a plate type heat exchanger. It is operated by an electro hydraulic servo valve mechanism. The following are features of the ACOC;

• Fail safe position is valve open for maximum cooling.

• Fire seal forms an air tight seal between the unit outlet and the cowling orifices.

• Control by either channel A or B of EEC.

• Valve position feed back signal via LVDT to each channel of EEC.

• Valve positioned by fuel servo pressure acting on a control piston.

• Fuel servo pressure directed by the electro hydraulic servo valve assembly which incorporates a torque motor.

The valve is operated via signals from the EEC heat management system. The electro-hydraulic servo valve directs a controlling fuel pressure to the operating piston. Depending on which side of the piston the fluid is present depends whether the valve opens or closes. An LVDT gives positional feedback to the EEC of the valves position.

Revision 2 Page 9-3


Revision 2 Page 9-4

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Heat Management System Air Control Valve Electro Hydraulic Servo Valve (EHSV) Purpose To provide the 'muscle' to move the air control valve to the EEC commanded position. Location Bolted to the air control valve casing. Description Two stage directional flow valve. The stages are;

• Stage 1 is an electrically activated torque motor and 'Jet pipe'.

• Stage 2 is a spool valve. The following are features of the EHSV;

• Two independent torque motor windings - one connected to each channel of EEC.

• Operation, from either channel of EEC.

• Jet pipe protected by 90 micron filter.

• Biased to ensure air control valve fully open at engine start condition.

• Single fuel servo supply from fuel metering unit (FMU).

Revision 2 Page 9-5


Revision 2 Page 9-6

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Heat Management System Fuel Diverter and Back to Tank Valve Purpose The fuel diverter valve and the back to tank valve together form a single unit. Command signals of the EEC control the two valves. The two valves in turn manage the flow of high pressure fuel. This is done to optimise the heat exchange process that takes place between the fuel and oil. Location The unit is bolted to the rear of the fuel/oil heat exchanger. Description Fuel Diverter Valve This valve is a two position valve and is operated by a dual coil solenoid. The control signals to energise/de-energise the solenoid come from the EEC

• Solenoid energised - mode 1 or 3.

• Solenoid de-energised (fail safe position) mode 4 or 5. Back to Tank Valve This valve is a modulating valve and will divert a proportion of the LP fuel back to the aircraft tanks as controlled by the EEC. The interface between the EEC and the valve is a modulating torque motor; the torque motor will direct HP servo fuel to position the valve.

• Fail safe position is with the valve fully closed - no fuel return to tank.

Revision 2 Page 9-7


Revision 2 Page 9-8

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Heat Management System Fuel Diverter and Return Valve Removal/Installation The removal/installation details for the fuel diverter and return valve is shown below. Reference must be made to the Engine Maintenance Manual. Removal CH 73-13-42-000-010. Installation CH 73-13-42-400-010. Warning: Be careful when you work on the engine components immediately after the engine is shutdown. the engine components can stay hot for up to one hour. Do not let engine fuel stay on your skin for a long time. flush the fuel from your skin with water. the fuel is poisonous and can go through your skin and into your body. Be careful during removal/installation of the fuel diverter and return valve because it weighs 13 lb. (6 kg). Caution: Do not let engine fuel or oil fall onto the engine. unwanted fuel or oil must be removed immediately with a clean lint free cloth. the fuel or oil can cause damage to the surface protection and to some parts. Do not put fuel that has been drained from the engine back in to the fuel system.

Removal • Drain the fuel from the fuel cooled oil cooler by

removing the fuel drain plug into a suitable container.

• Remove the FCOC from the engine with the FDRV attached to the FCOC.

• Remove the FDRV from the FCOC. Installation • Discard all used gaskets and packing and replace with

new items on reassembly.

• Install the FDRV to the FCOC.

• Torque the bolts with limits quoted in the Engine Maintenance Manual.

• Install the FCOC to the engine.

• Do a functional test of the fuel recirculation cooling system (73-13-42-720-010).

Revision 2 Page 9-9



© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Heat Management System Heat Management System Operation The following are the four modes of control for the heat management system. The system is fully automatic as controlled by the EEC. The four modes will be in effect when certain aircraft/engine operating conditions exist. Mode 1 This is the normal mode and is shown below. In this mode all the heat from the engine oil system and the IDG oil system is absorbed by the LP fuel flows. Some of the fuel is returned to the aircraft tanks where the heat is absorbed or dissipated within the tank. This mode is maintained if the following conditions are satisfied: -

• Engine not at high power setting (Take Off and early part of climb (not below 25,000ft).

• Cooling spill fuel temperature less than 100 deg C.

• Fuel temperature at pump inlet less than 54 deg C.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Heat Management System Heat Management System Mode 3 Mode 3 shown below is the mode that is adopted when the requirements for fuel spill back to tank can no longer be satisfied i.e.

• Engine at high power setting (below 25,000ft).

• Spill fuel temperature above limits (100 deg C).

• Tank fuel temperature above limits (54 deg C). In this condition the burned fuel absorbs all the heat from the engine and I.D.G. oil systems. If however, the fuel flow is too low to provide adequate cooling the engine oil will be pre-cooled in the air/oil heat exchanger, by a modulated air flow, before passing to the fuel/oil heat exchanger. This is the preferred mode of operation, when return to tank is not allowed




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Heat Management System Heat Management System Mode 4 Mode 4 is the mode adopted when the burned fuel flow is low. For example; Low engine speeds. High HP fuel pump inlet temperature. In this mode the fuel/oil heat exchanger is operating as a fuel cooler. The excessive heat is passed to the engine oil. The ACOC extracts the heat from the oil that has been heated up by the hot fuel.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Heat Management System Heat Management System Mode 5 Mode 5 is the mode that is used when the system conditions demand operation as in mode 3, but is not permitted due to;

• IDG oil system temperature is excessive.

• The fuel spill to the aircraft tank is not permitted because of high spill fuel temperatures.

Mode 5 is the adopted position for the fail safe engine conditions.






SECTION 10

COMPRESSOR AIRFLOW CONTROL (Chapter 75)

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Compressor Airflow Control System

Introduction Description/operation The engine incorporates two air bleed systems and a variable stator vane (V.S.V.) system, which together are used to: -

• Ensure stable airflow through the compressor at "off design" conditions.

• Ensure smooth, surge free, accelerations and decelerations (transient conditions).

• Improve engine-starting characteristics.

• In re-stabilising the engine if surge occurs (surge recovery).

The complete system comprises three sub-systems, which are: -

• An L.P. compressor air bleed located at engine station 2.5 and known as the Booster Stage Bleed Valve (B.S.B.V.)

• H.P. compressor air bleeds on stages 7 and 10.

• The V.S.V. system which comprises variable inlet guide vanes, at the inlet to the H.P. compressor, and 4 stages of variable stator vanes on the A1 and 3 stages on A5 engines.

The EEC controls all three systems A schematic overview of the complete airflow control system is shown below.


© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control


© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control L.P. Compressor Bleed Valve (LPCBV) - A5

Purpose The LPCBV bleeds air from the rear of the L.P. compressor at engine station 2.5, the bleed air is vented into the fan air duct. The bleed valve provides improved an improved surge margin during starting, low power and transient operations. The bleed valve is controlled by the EEC and is fully modulating, between the fully open and fully closed positions this is a function of: -

• N1 corrected speed

• Altitude

• Aircraft forward speed (Mach Number) For starting the bleed valve is fully open and will progressively close during engine acceleration, during cruise and take off the valve is fully closed. For decelerations and engine operation in reverse thrust the valve is opened. In the event of an engine surge the valve is opened to enhance recovery.

LPCBV Mechanical Arrangement The L.P. Compressor Bleed Valve is a continuous ring type valve that rotates and slides forward to open and rearward to close. Ten support arms support the ring. Two of the support arms are driven via a lever and actuating rod by both the LPCBV master actuator and the slave actuator. The two actuators utilise H.P. fuel pressure (from the FMU) as the hydraulic medium and are hydraulically linked to ensure simultaneous movement. The master actuator interfaces with the EEC via a torque motor control and LVDT feedback. The mechanical linkage is shown below.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control L.P. Compressor Bleed Valve (LPCBV) – A1

Purpose The LPCBV bleeds air from the rear of the L.P. compressor at engine station 2.5, the bleed air is vented into the fan air duct. The bleed valve provides improved an improved surge margin during starting, low power and transient operations. The bleed valve is controlled by the EEC and is fully modulating, between the fully open and fully closed positions this is a function of: -

• N1 corrected speed

• Altitude

• Aircraft forward speed (Mn) For starting the bleed valve is fully open and will progressively close during engine acceleration, during cruise and take off the valve is fully closed. For decelerations and engine operation in reverse thrust the valve is opened. In the event of an engine surge the valve is opened to enhance recovery.

LPCBV Mechanical Arrangement The annular bleed valve comprises 27 flaps that are attached by 25 link arms and 2 power arms to a synchronous ring. Two actuating rods connect the two power arms to two actuators. The two actuators utilise H.P. fuel as the hydraulic medium, and are hydraulically ‘linked’ to ensure simultaneous movement. The mechanical arrangement is shown below.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Booster Stage Bleed Valve

Component Description Actuators The two B.S.B.V. actuators utilise H.P. fuel as a hydraulic operating medium. The actuators are located on the rear of the intermediate casing on either side of the H.P. compressor, as shown below. Only one of the actuators, the one on the left hand side, interfaces with the EEC This actuator is called the Master actuator, the right hand side actuator is called the Slave actuator. The two actuators are hydraulically linked by two tubes, which pass across the top of the H.P. compressor case. The master actuator incorporates a Linear Variable Differential Transducer (L.V.D.T.) which transmits actuator positional information back to the EEC The slave actuator incorporates two overload relief valves, which prevent overpressurisation of the actuators in the case of faults, such as a mechanically seized actuator.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control BSBV Master Actuator Removal/Installation Removal/installation of the master actuator is quite straightforward. The disconnection points are shown on the diagram below. The following points should be noted: -

• All sealing rings must be discarded on removal and new sealing rings fitted on installation.

• All threads should be lubricated with clean engine oil on installation.

• Observe the torque loading quoted in the maintenance manual.

• The bolt, which secures the actuator fork end to the actuating rod, is "locked" by a double key washer. A new washer must be used on installation.

• Upon completion of the actuator change, carry out Test No 1 or 3 - leak test, followed by Test No 11 - High Power Assurance test.

The full procedure to remove/install the B.S.B.V. master actuator can be found in the aircraft maintenance manual.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control B.S.B.V. Slave Actuator - Removal/Installation Removal/installation of the slave actuator is quite straightforward. The disconnect points are shown below. The following points should be noted: -

• All sealing rings must be discarded on removal and new sealing rings fitted on installation.

• All threads should be lubricated with clean engine oil on installation.

• Observe the torque loading quoted in the maintenance manual.

• The bolt that secures the actuator fork end to the actuating rod is "locked" by a double key washer - a new washer must be used on installation.

Upon completion of the actuator change, carry out Test No 1 or 3 - leak test, followed by Test No 11 - High Power Assurance test. (Refer to section 14 page 14-31) For full removal/installation procedures refer to the aircraft maintenance manual CH 75-31-43.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Variable Stator Vane System (VSV) A1

Introduction Variable Incidence Stator Vanes control the entry of air into the H.P. compressor. The variable vanes control the angle at which the air enters the first five stages of the H.P. compressor. The angle varies with the H.P. compressor speed (N2); this reduces the risk of blade stall and compressor surge. The five stages of variable incidence stators comprise inlet guide vanes to stage 3 and stages 3, 4, 5 and 6 stator vanes.

Mechanical Arrangement Each vane has pivots at its inner and outer ends, which allow the vane to rotate about its longitudinal axis. The outer end of each vane is formed into a shaft which passes through the compressor case and is attached by a short lever to a 'unison ring', (one unison ring for each stage). Short rods to a crankshaft connect the five unison rings. A short rod to an actuator that utilises H.P. fuel as a hydraulic operating medium connects the crankshaft. Signals from the EEC direct H.P. fuel to extend/retract the actuator. Actuator movement causes the crankshaft to rotate, and, through the unison rings, reposition the variable stator vanes.

The actuator incorporates an LVDT which signals actuator positional information back to the EEC.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Variable Stator Vane System (VSV) A5

Introduction Variable Incidence Stator Vanes control the entry of air into the H.P. compressor. The variable vanes control the angle at which the air enters the first five stages of the H.P. compressor. The angle varies with the H.P. compressor speed (N2); this reduces the risk of blade stall and compressor surge. The four stages of variable incidence stators comprise inlet guide vanes to stage 3 and stage 3, 4, and 5 stator vanes.

Mechanical Arrangement Each vane has pivots at its inner and outer ends, which allow the vane to rotate about its longitudinal axis. The outer end of each vane is formed into a shaft, which passes through the compressor case and is attached by a short lever to a Unison ring, (one unison ring for each stage). Short rods to a crankshaft connect the four unison rings. A short rod to an actuator that utilises H.P. fuel as a hydraulic operating medium connects the crankshaft. Signals from the EEC direct H.P. fuel to extend/retract the actuator. Actuator movement causes the crankshaft to rotate, and, through the unison rings, reposition the variable stator vanes.

The actuator incorporates an L.V.D.T. which signals actuator positional information back to the EEC




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Variable Stator Vane System (V.S.V.) Actuator Removal/Installation Prior to VSV actuator removal, it is necessary to drain the fuel lines to the actuator. The fuel lines are drained at the union locations shown below. Fuel is drained at this point because it is the lowest point in the system and also because fuel drained here is less likely to cause contamination of the engine electrical harness.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Variable Stator Vane System (V.S.V.) Actuator Rigging The full procedure is in Chapter 75-32-41 of the engine maintenance manual but is summarised below.

• After refitting the unit, rig pin the actuator piston in the high speed position (there is only one rig pin position).

Note: This is achieved by moving the ram to the fully retracted position against the high speed stop, then withdrawing the ram as necessary to align the rig pin hole in the fork end with the hole in the rig pin housing (Detail B).

• Rig pin the VSV crankshaft in the high speed position.

• Connect the rod adjusting its length to suit (it has a left and right hand threads – ‘turnbuckle effect’). Ensure control rod ends are in ‘safety’ on completion.

• Remove rig pins.

• Upon completion carry out Test 3 or 1 leak checks, followed by Test No 11 - high power assurance test.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Variable Stator Vane System (V.S.V.) Actuator Removal/Installation The actuator’s disconnect and mounting features are shown below. A full description of the removal/installation can be found in Chapter 75-32-41 of the Maintenance Manual. Points to Note:

• Rig pin the VSV crankshaft before disconnecting the actuator see next illustration.

• Drain off fuel from the lowest point to avoid contaminating the engine electrical harness.

• Installation is the reverse of the removal procedure.

• See next page for rigging.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Handling Bleed Valves Introduction Handling bleed valves are fitted to the H.P. compressor to improve engine starting, and prevent engine surge when the compressor is operating at off-design conditions. A total of four bleed valves are used, three on stage 7 and one on stage 10. 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 as a function of N2 and T2.6 (N2 corrected). When the bleed valves are open, H.P. compressor air bleeds into the fan duct through ports in the inner barrel of the 'C' ducts. The servo air used to operate the bleed valves is H.P. compressor delivery air known as P3 or Pb. The bleed valves are arranged radially around the H.P. compressor case as shown below. Silencers are used on some bleed valves. All the bleed valves are spring loaded to the open position and as a result will always be in the correct position (open) for starting.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Handling Bleed Valves - Location The diagram below shows the location of the four bleed valves and the solenoid control valve.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Handling Bleed Valves - Operating Schedule The handling bleed valves have three operating regimes:

• Steady State.

• Transient - Acceleration/Deceleration.

• Surge/Reverse Operation of the bleed valve is scheduled against N2 corrected for changes of H.P. compressor (T2.6) inlet temperature - known as N2C26.

Steady State The valves are commanded ‘open’ whenever N2C26 is below the steady state closing speed.

Transient The valves are commanded ‘open’ at the beginning of accelerations/decelerations and will ‘close’ when either the speed limits are exceeded or timers expire. After an acceleration phase has ceased, the valve will remain open until a period of 5 seconds have elapsed, after which it will then be signalled to close. During a deceleration phase the valve will remain open until a period of 62 seconds have elapsed after the engine has stabilised at the new engine speed.

Surge/Reverse The valves will be commanded open in the event of a surge. In reverse thrust laws similar to the transient laws apply.



BLEED SCHEDULES FOR V2500-A5/D5

Condition N2C26 7A status 7B status 7C status 10 status SS ONLY SS ONLY SS & TR SS ONLY

starting <8623 open open - closes before

reaching idle open - closes on

reaching idle open - closes before

reaching idle idle/taxi 8623 open closed closed* closed

take off acceleration 8623<N2C26<12100 open>closed closed

opens on detection of acceleration, then closes at mid-power

closed

take off (including derates)

Begin T/O 12100, 90% de-rated 12044, 80%

derated 11965 closed closed closed* closed

climb Begin 11869, Mid

12142, End 12294 closed closed closed* closed

cruise (mid) ~12100 closed closed closed* closed

end of cruise deceleration 12000<N2C26<10819 closed>open closed

opens on detection of deceleration, then

closes closed

top of descent 10819 open closed closed* closed mid descent 10211 open closed closed* closed end of descent 8509 open closed closed* closed approach 9085<N2C26<11560 open closed closed*,*** closed touchdown 9745 open closed closed* closed

reverse 12135 open (if N2C26 below

certain threshold) closed open (if N2C26 below certain

threshold) closed

idle/taxi <8623 open closed closed* closed

surge recovery NA open (if N2C26 below

certain threshold) closed open (if N2C26 below certain

threshold)

open (if N2C26 below certain threshold)

* bleed valve will open in response to throttle lever angle variation

** the holding condition varies based on aircraft weight, landing runway altitude, airport traffic, typical mission etc. the EEC does not have a unique TRA position for holding conditions. generally a 30% maxmum take off thrust is used for holding condition power setting.

*** bleed valve will open when approach mode is selected and engine switches from low to high idle


© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Handling Bleed Valves - Operation The bleed valves and the solenoid control valves all operate in the same manner. The operation of one bleed valve only is described.

Bleed Valves The 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 bleed valves will be in the correct position - open - for engine start.

When the engine is started the bleed air will try to close the valve. The valve is kept in the open position by servo air (P3) supplied from the solenoid control valve, (solenoid de-energised) as shown below.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Handling Bleed Valves Operation The EEC will close the bleed valves at the correct time during acceleration. The bleed valve is closed by the EEC, which energises the solenoid control valve, as shown below. Energising the solenoid control valve vents the P3 servo air from the opening chamber of the bleed valve, and the valve will move to the closed position.

During an engine deceleration the reverse operation occurs and the bleed valve opens.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Handling Bleed Valves – Troubleshooting. The HP Compressor rotor path lining is designed to be worn away by the rotor blades of the compressor. As air is expelled through the bleed valves during their normal operation, some of the debris from the rotor path lining contaminates the valve seals. If this lining becomes lodged in the carbon seals within the bleed valve then there is the possibility that this will prevent the valve from operating smoothly and the valve will seize. If the valve does not operate when required, the engine will experience problems at critical points. Handling Bleed valves failures have two major consequences:

• Hung Start: If the bleed valve sticks in the closed condition, (Non-detected FADEC fault) or the solenoid valve sticking in the ‘energised position (Non-detected FADEC fault) the engine will experience difficulty in starting ‘hung start or surge’ during the start cycle. This is due to the fact that during low engine speeds, the bleed valve system is designed to effectively ‘dump’ into the ‘C’ duct a large amount of the air supplied to the HP Compressor. This is necessary because of the HP Compressor’s inability to handle all the mass flow of air being supplied to it by the LP Compressor during low speed operation.

All bleed valves (3 off stage 7, 1 off stage 10) are spring loaded to the ‘open’ position for engine starting. They are held open during engine running, by solenoid valve directed P3 air. According to schedule requirements, the bleed valves will close progressively during the starting cycle in the sequence 7B, 10, and 7C. The 7A valve stays open up to and above idle.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Compressor Airflow Control Handling Bleed Valves – Troubleshooting Continued: If a bleed valve fails to close when required to do so, under certain conditions the engine may exceed the recommended EGT operating limits thus preventing the aircraft from taking off. This will be caused as a result of the EEC trying to achieve Take-off EPR but with a reduced volume of air being supplied to the combustion chamber for mixing with fuel, ignition and subsequent expansion. Therefore the EEC makes up for the shortfall in the available volume of air and simply demands the FMU to provide more fuel to compensate. The resultant ‘over-fuelling’ provides the required EPR, but with the penalty of increased EGT. TSM Supporting data 75-00-00-301 (Bleed Valve Troubleshooting Hints) give comprehensive recommendations in the diagnosis of bleed valve related problems. Rigorous troubleshooting would reduce large number of NFF cases. The lubrication of bleed valves should not be carried out.

Engine Parameter shift/mismatch during climb/cruise Engine Parameter shifts due to an open bleed valve that are not noticed at engine start are more likely to become evident at higher EPR power settings. This increases the likelihood of an EGT Exceedance/Overlimit. Additionally, an open bleed valve as a result of bleed valve system problems will also result in unexplained engine parameter shifts. Possible causes:

• Bleed valve(s) stuck open (Non-detected FADEC fault).

• Solenoid valve sticking in the de-energised position (Non-detected FADEC fault).

• An electrical failure of the solenoid valve which results in the solenoid moving to the de-energised position (FADEC fault).




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Compressor Airflow Control Compressor Airflow Control Handling Bleed Valves – Troubleshooting Continued: Engine Operation Impact (Transient Manoeuvres and Surge Recovery).

Engine Operation Impact During transient manoeuvres (acceleration/deceleration) and surge recovery, HP compressor stability is maintained by opening particular bleed valves as defined by the EEC logic. For transient manoeuvres on the ground, the 7C are opened and during flight both the 7A and 7C are opened. In both cases the 7C are opened based upon a transient detect and is closed after a set period of time has elapsed. There are no valves opened at take-off power or steady state cruise. For surge recovery, the 7A, 7C and 10 stage bleed valves are opened to maintain compressor stability.

Possible causes Engine problems (stall/surge) on transient operation can be the result of:

• Bleed valve(s) not being opened during the transient (acceleration/deceleration)

• Bleed valve closing early.

• Bleed valves not being open can be due to the bleed valve sticking in the closed position (Non-detected FADEC fault), or the solenoid sticking in the energised position (Non-detected FADEC fault).

• Bleed valves closing early can be due to the solenoid sticking such that the full de-energised position is not obtained (Non-detected FADEC fault), bleed valve internal seal wear and leakage of P3 servo air.


SECTION 11

ENGINE SECONDARY AIR SYSTEMS A.C.C System (Chapter 75)

Make up Air System (Chapter 75) A.C.A.C System (Chapter 75)

Aircraft Services Bleed (Chapter 36)

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Secondary Air Systems Engine Secondary Air systems Introduction Purpose The secondary air systems serve the function of the engine and aircraft. Description The secondary air system is made up of the following:

• Active clearance control system (ACC).

• 10th stage make up air system.

• Aircraft services bleed system.

• Air cooled air cooler (ACAC) for the no.4 bearing cooling and sealing.

Active Clearance Control (ACC) The system improves engine performance by ensuring that the HPT and LPT operate with optimum turbine blade tip clearances. This is achieved by directing a controlled flow of cooling air to reduce the thermal growth of the turbine casings. This minimises the increase in turbine blade tip clearances which otherwise occurs during the climb and cruise phases. 10th Stage Make Up Air System The purpose of this system is to provide additional cooling airflow to the HPT stage 2 disc and blades. The cooling air used is taken from the 10th stage manifold and is controlled by a two position pneumatically operated valve.

The valve position is controlled by the EEC as a function of corrected N2 and altitude. Aircraft Services Air Offtake System The engine supplies the aircraft with bleed air taken from HPC stages 7 and 10. The air that is taken from the engine is used for the following:

• Cabin pressurisation and conditioning.

• Wing anti icing.

• Engine cross feed starting.

• Hydraulic system pressurisation.

• Water system pressurisation. The required air is bled from the HPC of each engine. Air cooled air cooler (ACAC) HPC12 air is used for cooling and sealing the no.4 bearing in the centre bearing compartment. The ACAC pre cools the HPC12 air prior to the air being passed to the centre bearing compartment. The cooled HPC12 air is commonly known as buffer air. The ACAC uses fan bypass air as the cooling medium.


© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Secondary Air Systems


SECONDARY AIR SYSTEMS INTRODUCTION DET

V250

280 HPC STAGE 7 AIR OFFTAKE HPC STAGE 10 AIR OFFTAKE

MAKE UP AIR VALVE

AIR COOLED AIR COOLER (ACAC)

ACC ACTUATOR

MODULATING AIR CONTROL VALVE ACC TUBES FOR THE

LPT AND HPT

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Secondary Air Systems Active Clearance Control (ACC) Purpose The system improves engine performance by ensuring that the HPT and LPT operate with optimum turbine blade tip clearances. This is achieved by directing a controlled flow of cooling air to reduce the thermal growth of the turbine casings. This minimises the increase in turbine blade tip clearances which otherwise occurs during the climb and cruise phases. Location The ACC system is mainly located about the core engine. Description The ACC system consists of the following items:

• LPT and HPT cooling manifolds.

• Operating actuator with LVDT feedback.

• Modulating air control valve unit.

• EEC control.

• Fan bypass air cooling medium. The EEC controls the ACC system by monitoring the following parameters:

• Corrected N2.

• Aircraft altitude. From these two parameters the EEC will signal the operating actuator.

The operating actuator moves in a linear motion by the influence of fuel pressure. The EEC receives feedback of the actuator position by an LVDT. The operating actuator moves a linkage that controls the valves in the modulating air control unit for the LPT and HPT case cooling. The HPT and LPT casings are cooled by fan bypass air that is ducted from the fan bypass. Failsafe Position Upon the event of fuel pressure loss and/or EEC power failure the ACC modulating air control valves will adopt the following positions;

• LPT is –44% open.

• HPT is closed.





Page 11-5 Revision 2

At engine shut down or nil servo pressure the ACC actuator will assume the failsafe position. A spring in the ACC actuator will force the piston to the failsafe position.

Active Clearance Control Components ACC Actuator Purpose The ACC actuator provides the movement to the modulating air valves so it can vary the LPT and HPT cooling airflows. Location The ACC actuator is located on the right hand side of the core engine in the 5 o’clock position. It is mounted on the compressor casing. Description The ACC actuator consists of the following: • Linear motion two directional piston. • Dual track LVDT. • Electro hydraulic torque motor. • Filter. The ACC actuator receives signals from the EEC. The torque motor will direct high pressure fuel to one of the two sides of the piston. This is dependent on the EEC command signal. Piston movement will result in a movement in the push pull rod that links the ACC actuator and the modulating air valve. The LVDT will feedback the piston position to the EEC.

Modulating Air Control Valve Unit Purpose The modulating air control valve receives ducted air from the fan bypass stream and regulates, as per ACC actuator input, the flow rate to the LPT and HPT ACC manifolds. Location The modulating air control valve is located on the right hand side of the core engine in the 5 o’clock position. It is mounted on the turbine casing. Description The modulating air control valve has two separate valves. They are: • HPT valve. • LPT valve. The two valves are designed to operate to allow the optimum airflow to the respective casings. The failsafe position is: • HPT is closed. • LPT is –44% open.


Page 11-6

ACC ACTUATOR MODULATING AIR CONTROL VALVE

Revision 2

DET

V250

282

ACC ACTUATOR AND MODULATING AIR VALVE

FAN BYPASS AIR INLET

TORQUE MOTOR

LVDT FEEDBACK

SERVO FUEL SUPPLIES SERVO FUEL SUPPLIES

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Secondary Air Systems Active Clearance Control Components HPT ACC Manifold Description The HPT ACC manifold is designed to impinge cooling air onto the turbine casing about the rotor blade path. This is done to reduce the rotor blade tip to rotor path gap. Location The HPT ACC manifold is located on the HPT casing. Description The assembly consists of left and right hand tube assemblies, which are a simple push fit into the manifold. The tube assemblies are sealed off at their upper ends. Air from the air control valve enters the manifold and is directed to the left and right tubes. Air outlet holes on the inner face of the tubes direct the air onto the HPT casings.

LPT ACC Manifold Description The LPT ACC manifold is designed to impinge cooling air onto the turbine casing about the rotor blade path. This is done to reduce the rotor blade tip to rotor path gap. Location The LPT ACC manifold is located on the LPT casing. Description The assembly consists of upper and lower tube assemblies with integral manifolds; both ends of the cooling tubes are sealed. Air from the air control valve enters a supply tube, which then splits to feed air into two tubes that supply the upper and lower manifolds. The manifolds direct the air into the cooling air tubes. Air outlet holes on the inner surfaces direct the air onto the LPT cases.




ACC LPT MANIFOLD HALVES

ACC HPT MANIFOLD HALVES

DET

V250

283

HPT AND LPT DISTRIBUTION MANIFOLDS



The dual track LVDTs will send feedback signals to the EEC of the ACC system operation.

ACC System Operation The operation of the ACC system is as follows: The EEC controls the opening and closing of the ACC system by monitoring input signals of; • Corrected N2. • Altitude. The EEC commands an input signal to the torque motor. The torque motor positions the jet pipe servo valve. The torque motor can deflect the jet pipe servo valve to bias the direction of flow of the servo fuel pressure. The jet pipe servo valve controls the direction of flow of servo fuel pressure to effectively move the pilot valve. The pilot valve moves and admits servo fuel pressure to either side of the piston. Servo fuel pressure will act on one side of the piston at any one time when a movement is required. The movement of the piston moves a push pull rod that in turn operates the modulating air control valve. When stabilisation of the piston is required the EEC will cancel the input signal to the torque motor. This allows the jet pipe to return to the central position and as a result of this the pilot valve will move into the equilibrium position. Servo fuel pressure is now present on both sides of the pilot valve. The spring will bias the pilot valve position by forcing it to one side.

ACC Operating Schedule The graph represents the conditions of engine operation and the effect it has on the modulating air valves position. Position A At position A the engine is shut down. This is also the failsafe position. HPT ACC valve is closed. LPT ACC valve is at –44%. Position B This position represents idling conditions. HPT ACC is closed. LPT ACC is closed. Position C This position represents a typical take off condition. This position is altitude dependent. HPT ACC is starting to open. LPT ACC is at 70%. Position D and E These positions represent typically cruise and top of descent conditions. This position is altitude dependent. HPT ACC at D is 30% and at E is fully open. LPT ACC is fully open at points D and E.



© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Secondary Air Systems 10th Stage Make Up Air System Purpose The 10th stage make up air valve system allows additional cooling air to the HPT stage 2 disc and blades. Description The 10th stage make up air valve system consists of the following components:

• EEC control.

• Make up valve control solenoid.

• Two position type on/off valve.

• Microswitch positional feedback. The EEC uses input signals of:

• Corrected N2.

• Altitude. The EEC to signal the control solenoid to operate uses these inputs. The control solenoid manages the flow of P3 (HPC stage 12) air for the pneumatic operating medium. The two position make up air valve either opens to flow stage 10 air or closes for no flow. The solenoid is de energised when the valve is in the open position. This is the fail safe position.

The microswitch gives a positional feedback signal to the EEC indicating of either:

• Valve open.

• Valve closed. The valve is open for all conditions of flight/engine operation except for cruise. In cruise the valve is closed.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Secondary Air Systems 10TH Stage Make Up Air System Components 10th Stage Make Up Valve Purpose The 10th stage make up valve purpose is to supply air to supplement the normal airflows around the no.4 bearing housing and the HPT disc and blades. Location The 10th stage make up air valve is located at the top of the HPC casing. Description The 10th stage make up air valve consists of the following components:

• Operating piston.

• Microswitch feedback.

• Valve body. The valve is a two positional type. It can either allow flow HPC stage 10 air or cut it off from the engine. There is no modulation. The operating piston is spring loaded to the open position when servo air is not present in the piston chamber. Servo air is used to close the valve. The micro switch gives positional feedback of the piston position hence the valves condition. The fail safe position is valve open.

Control Solenoid valve Purpose The control solenoid valve purpose is to manage the flow of servo air pressure to the make up air valve. Location The control solenoid is located on the right hand side of the fan case approximately in the 4 o’clock position. Description The control solenoid consists of the following components: Solenoid pack. Pilot valve. Valve body. The solenoid control valve will direct the flow of servo air pressure to port when it is de-energised. The solenoid control valve will direct the flow of servo air pressure to the make up valve when it is energised.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Secondary Air Systems 10TH Stage Make Up Air System Operation The operation of the system is as follows: The EEC constantly monitors:

• Corrected N2.

• Pressure altitude. The make up air valve will be commanded to open during all flight conditions except during cruise. To Energise the Solenoid Valve The solenoid pack attracts the cover plate A towards it thus opening up the chamber that is at the spring side of the pilot valve. The servo air pressure and spring pressure at the spring end of the pilot valve overcomes the servo air pressure alone on the opposite side of the pilot valve. This makes the pilot valve move towards the cover plate B. Cover plate B is pushed away from the orifice allowing servo air to enter the make up valve. The servo air enters the make up valve piston chamber in the opposite side to the spring side. The servo air pressure overcomes the spring pressure and forces the piston to move and hence close the valve. The microswitch contacts are broken and a feedback signal is fed to the EEC.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Secondary Air Systems To De-Energise the Solenoid Valve The solenoid pack is de-energised. The cover plate A is force away from the solenoid pack by spring pressure and forced against the orifice. This prevents servo air from entering the spring side of the pilot valve. Servo air pressure on the opposite side of the pilot valve now forces this valve against the spring. As the pilot valve moves cover plate B closes off the orifice. This prevents servo air pressure from entering the make up valve piston chamber. The spring affecting the piston valve forces the piston to move and open the make up valve orifice. The microswitch contacts are made and a feedback signal is fed to the EEC. Fail Safe Position The fail safe position has the valve in the open position hence solenoid de-energised. This is also true if a power loss is experienced or a loss of servo air pressure.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Secondary Air Systems Aircraft Services Air Offtake System Purpose To provide the following aircraft systems with engine ducted air supply:

• Cabin pressurisation and conditioning.

• Wing leading edge anti icing.

• Engine cross bleed starting.

• Hydraulic system pressurisation. Location The bleed air offtakes are taken from:

• HPC stage 7 for high power conditions.

• HPC stage 10 for low power conditions. Description HPC air is taken from the engine and ducted towards the aircraft services. The HPC stage 7 offtake has a non return valve installed before the two offtakes join. The NRV protects against HPC stage 10 air from reverse flowing back into the HPC stage 7 of the engine. The HPC stage 10 offtake has a control valve called the high pressure valve (HPV). After the two offtakes come together as one there is a pressure regulating valve (PRV). A switch located in the flight deck controls the PRV.

The over pressurisation valve (OPV) protects the system against excessive pressures. The precooler prepares the bleed air to an acceptable temperature before it enters the environmental control system (ECS). The pre cooler utilises fan bypass air to cool the HPC bleed air. The temperature limiting thermostat (TLT) controls the PRV when an over temperature has been experienced. The temperature controlling thermostat (TCT) controls the pre cooler valve but if bleed temperature cannot be maintained the temperature limiting thermostat will signal for the PRV to close. The bleed monitoring computer controls the functions of bleed air system making the system fully automatic.






The PRV is spring loaded closed when there is no pneumatic air available.

Aircraft Services Air Offtake System Operation The following is the operation of the aircraft services air offtake system. The bleed monitoring computer (BMC) controls the opening and closing of the;

• Over pressure valve (OPV).

• Pressure regulating valve (PRV) a selector switch in the flight deck also controls (this.

• Fan air valve (FAV).

• High pressure valve (HPV). The bleed monitoring computer monitors the following:

• Temperature limiting thermostat (TLT).

• Temperature controlling thermostat (TCT).

• Pressure sensor downstream of the HPV.

• Pressure sensor downstream of the PRV. The selection of HPC stage 7 (also called IP bleed air) or HPC stage 10 is automatically done by the BMC. High pressure valve (HPV) The HPV will regulate HPC stage 10 air to 36+/- 3 psi. The HPV will close if upstream pressure is greater than 100 +/- 5 psi and/or downstream HPC stage 7 greater than 36 +/- 3 psi. Pressure regulating valve (PRV)

The PRV will regulate airflow to 44 +/- 3 psi. The PRV starts to open at approximately 8 psi. Over pressure valve (OPV) The OPV will start to close at 75 psi. (Depending on mod standard 79 psi) The OPV will be fully closed at 85 psi. The OPV will reopen at 35 psi. (Depending on mod standard 20 - 57 psi) Temperature limiting thermostat (TLT) The TLT will start to close the PRV to reduce pressure at 235 deg.c. The TLT over temperature is 247 deg.c. Above this value will reduce PRV pressure to 17.5 psi. The TLT maximum temperature is 257 +/- 3 deg.c. (60 sec delay) Above this value and the PRV is closed. Temperature controlling thermostat (TCT) TCT will regulate the temperature of the air entering the aircraft system to 200 +/- 15 deg.c. The TCT controls the opening/closing of the Fan air valve to regulate the fan airflow through the Pre – cooler. The pressure sensors feedback pressure signals to the BMC.



© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Secondary Air Systems Air Cooled Air Cooler (ACAC) Purpose The ACAC purpose is to pre cool HPC12 air. The ACAC uses fan bypass air as the cooling medium. Location The ACAC is located on the turbine casing. Bottom left hand side in the 5 o’clock position. Description The ACAC is a fin and tube type design. The fan bypass airflow that is utilised by the ACAC extracts heat from the HPC12 air. The HPC12 air is taken off the engine through a singular tube. The HPC12 air enters the ACAC and the heat exchange process takes place between the fan bypass air and the hot HPC12 air. The fan bypass air is ejected to atmosphere. The cooled HPC12 air leaves the ACAC and is distributed to the centre bearing compartment through three tubes. The tubes enter the diffuser casing in three positions. They are:

• 12 o’clock.

• 3 o’clock.

• 9 o’clock.

The cooled HPC12 (buffer) air enters the cooling jacket of the centre bearing chamber. The buffer air protects the no.4 bearing from excessive heat exposure. The buffer air enters the bearing compartment to prevent oil loss. This also pressurises the bearing chamber to allow the oil and air mix to leave the bearing chamber and enter the de oiler. The centre bearing compartment does not have oil scavenge pump.


© IAE International Aero Engines AG 2000


IAE V2500 Line and Base Maintenance Engine Secondary Air Systems

FAN BYPASS AIR INLET AIR COOLED AIR COOLER (ACAC)

HPC12 (BUFFER) AIR INTO THE CENTRE BEARING

COMPARTMENT

HPC12 AIR INTO ACAC

FAN BYPASS AIR OVERBOARD DUMP

COOLED HPC12 AIR OUT OF THE ACAC

DET

V250

376

CENTRE BEARING COMPARTMENT COOLING

JACKET AIR COOLED AIR COOLER (ACAC)

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Secondary Air Systems Active Clearance Control System HarnessThe HPT and LPT ACC system harness electrical connections are shown below.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Secondary Air Systems Miscellaneous Systems HarnessThe miscellaneous systems harness electrical systems are shown below.




SECTION 12

ENGINE ICE PROTECTION SYSTEM (Chapter 30)

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Ice Protection System Engine Ice Protection System Introduction Purpose Ice may form in the inlet cowl when the engine is operating in conditions of low temperature and high humidity. Ice build about the inlet cowl leading edge could affect engine performance and could cause engine damage from ice ingestion. To prevent ice formation ice protection systems have been incorporated into the engine. The inlet cowl leading edge is thermally ice protected. The P2/T2 probe mounted in the inlet cowl is thermally ice protected. The spinner of the fan module is ice protected by a flexible rubber tip. Description The engine ice protection system description is as follows: Inlet cowl ice protection The inlet cowl is thermally heated to prevent ice formation at the leading edge of the intake lip. The ice protection system for the inlet cowl is controlled from the flight deck by a selector switch. The switch will control the opening and closing of the TAI valve. The valve will allow the airflow taken from the HPC stage 7 to flow to the distribution manifold in the inlet cowl leading edge lip.

The distribution manifold will allow hot air to enter the inlet cowl leading edge. The excessive air is ejected overboard via an outlet located on the right hand side of the inlet cowl. Fault indications for the ice protection system are as follows: The flight deck anti icing selector switch illuminates. An ECAM warning message is generated. P2/T2 probe heater The P2/T2 probe is continuously heated during engine operation by an integral 115V heating coil. Spinner A solid rubber nose tip that vibrates naturally to break up and dislodge the ice immediately it starts to form protects the spinner against ice build up. Ground running Icing conditions may occur when the outside air temperature (OAT) is less than: 5.5 deg.c (42 deg.f). The humidity is high for example rain, sleet, snow, fog (visibility is less than one mile). If the above conditions exist the ice protection system must be operated as soon as the engine stabilises at low idle conditions after an engine start.


© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Ice Protection System


© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Ice Protection System Component Description Anti Icing Control Valve Purpose The anti icing control valve allows the flow of HPC stage 7 air to enter the TAI manifold in the intake cowl. Location The anti icing valve is located on the right hand side of the fan case in the 4 o’clock position. Description The anti icing control valve has the following function:

• On/off selection from the flight deck to allow the flow of warm air to the TAI manifold.

The anti icing valve is made up of the following items:

• Valve body.

• Linear moving piston.

• Control solenoid.

• Air filter.

• Butterfly valve.

• Micro switch.

• Manual override (as per MMEL requirements).


© IAE International Aero Engines AG 2000 Engine Ice Protection System IAE V2500 Line and Base Maintenance

Revision Page 12-4

ANTI ICE VALVE FILTER

LOCKOUT PIN

ELECTRICAL CONNECTOR

VALVE BODY

ENGINE ANTI ICE VALVE ENGINE ANTI ICE VALVE 2

DET

V250

273

DET

V250

273

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Ice Protection System Operation Valve closed The control solenoid is energised. The ball valve of the control solenoid is hard against the ambient vent outlet. This prevents upstream air from escaping to vent. The pressure acting on the piston at position area A is greater than the pressure acting on position area B. This along with spring pressure holds the butterfly valve in the closed position. Valve open The control solenoid is de energised. The ball valve is no longer held by the control solenoid against the ambient vent. The ball valve moves by spring pressure against the orifice, which allows upstream air to enter area A. This now prevents air passage to area A of the piston. The air pressure now remains at piston area B only. This pressure is greater than the spring pressure alone therefor the piston moves against the spring pressure. The resultant movement opens the butterfly valve and allows HPC stage 7 air to flow towards the TAI manifold. Fail safe position The fail safe position is as follows: • Solenoid de energised. • Servo air pressure at piston area B only. • Butterfly valve in the open position.

Manual override The valve has provision for being secured in either the; • Locked position. • Open position. This requirement is necessary when a valve has failed. The MMEL will advise of the actions required to allow despatch of the aircraft.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Ice Protection System ECAM Indications The engine anti icing valve has a micro switch, which will feedback the valve position in relation to the selector switch position. When the anti icing system is functioning normally a caption will appear:

• ENG A. ICE The caption appears on the upper ECAM lower right hand side. The selector switch will have the on indicator illuminating in the colour of blue. If a disagreement exists between the selector switch and the microswitch output signal to the EEC a fault has been detected. The fault detection occurs when one of the following situations exists:

• A valve failure to open.

• A valve failure to close. The fault portion of the selector switch will illuminate in the colour of amber when a disagreement exists. Th upper ECAM screen will display a WARNING and STATUS message of:

• ENG 1(2) VALVE CLSD.

• ENG 1(2) VALVE OPEN. These messages relate to the switch position and the intended valve position. The messages are engine specific.

The message is status related it therefore becomes a despatch critical message. Advice from the MMEL is required. Note: It is advisable not to lock the TAI valve in the open position for the higher thrust engines. If the valve fails in the closed position it is advisable to avoid icing conditions. If the valve fails in the open position there will be a thrust limit penalty. One or both may be inoperative provided the valve has failed in the open position and the performance penalties are applied and OAT does not exceed ISA +35 deg.c. Engine anti ice valve fault: For ER operations only one valve allowed to be failed in the closed position and providing the aircraft is not operating in icing conditions. Engine anti ice fault light: One or both valves may be inoperative provided the faulty valve is deactivated and considered inoperative in the open position.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Ice Protection System Ice Protection System Maintenance A failure of an anti icing valve to either open or close when commanded to do so by the flight deck selector switches, allows the engineer to override the valve in the open or closed position. The manifold off-take from the engine is mounted on the core engine and works its way towards the right hand side of the fan casing. At the fan casing an anti ice valve is located. There is a connection between the fan casing manifolds and the air intake manifolds. The air intake manifolds allows the HPC stage 7 air to flow into a distribution manifold. Excess air then under its own pressure is ejected to atmosphere from an outlet grid found on the right hand side of the intake cowl. The maintenance items that are to be discussed in this section are:

• TAI valve manual override; Deactivation AMM ref. 30-21-00-040-010. Reactivation AMM ref. 30-21-00-440-010.

• TAI valve change; Removal AMM ref. 30-21-51-000-010. Installation AMM ref. 30-21-51-400-010.

• Visual inspection of the anti icing supply ducts; Inspect AMM ref. 30-21-49-200-010.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Ice Protection System TAI Valve Manual Override As per MMEL requirement the anti icing valve can be manually overridden for continued engine operation. The following is the procedure for deactivation. Deactivation AMM ref. 30-21-00-040-010 • Carry out flight deck engine isolation procedures as per

AMM requirements.


• Prepare to lock the anti icing valve in the closed or open position. Turn the manual input shaft to lock the valve in the closed position.

The anti icing valve is spring loaded to the open position.

• When the valve is in the desired position insert the lock pin.

• Place a warning notice in the flight deck to inform of the valves condition.

• Put the engine to the normal operating condition. Reactivation AMM ref. 30-21-00-440-010 • Carry out the flight deck engine isolation procedures.

• Open the fan cowl doors.

• Remove the lockout pin and stow in the storage bracket provided.

• Put the engine to the normal operating condition.

• Inspect the fan cowl door area about the starter access/blow off door for de-lamination.

• Carry out the partial power assurance test for confirmation as to the anti icing system satisfactory operation.




ANTI ICE VALVE MANUAL OVERRIDE DET

V250

277

FADEC POWER SELECT SWITCH

ENGINE START CONTROL PANEL HAND TUR

POINT NING

HAND TURNING POINT

ANTI ICE VALVELOCKOUT PIN DEACTIVATION

HOLE

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Ice Protection System TAI Valve Change The following gives an outline for the procedure for the removal and installation of the TAI valve. Removal AMM ref. 30-21-51-000-010 • Carry out flight deck engine isolation procedures as per

AMM requirements.


• Disconnect the electrical connector and remove the coupling clamps.

• Loosen the duct at the intake cowl bulkhead.

• Remove the valve.

• Install blanks at to the valve openings. Install AMM ref. 30-21-51-400-010 • Remove blanks from the TAI valve.

• Install the valve and secure coupling clamps.

• Ensure that the direction of flow arrow is pointing in the correct direction.

• Tighten clamps to the AMM recommended torque values and install the electrical connector.

• Close the fan cowl doors.

• Return engine to its normal operating condition.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Ice Protection System Ice Protection System Inspection The ice protection system ducting and associated hardware requires inspection to ensure that the ice protection system is secure and free from defects. Visual Inspection of the Ice Protection System Ducts and associated Hardware. AMM ref. 30-21-49-200-010 • Carry out flight deck engine isolation procedures as per

AMM requirements.

• Open the fan cowl doors to allow access to the ice protection system.

• Carry out a full system detailed inspection noting for the following; Loose connections and fasteners. Cracks. Nicks. Tears. Galling. Pitting. Dents. Chafing.

The AMM has the accept/reject inspection standards that apply to the ice protection system components.






SECTION 13

ENGINE INDICATIONS (Chapter 77)

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications FADEC/Aircraft Interface Introduction Purpose The FADEC system supplies the aircraft systems with the relevant engine data in order to assist the aircraft to carry out its functions. Description The aircraft systems that it uses to interpret the engine data and display it for flight crew use is:

• Electronic centralised aircraft monitor (ECAM) system.

• Flight warning computer (FWC).

• Data management computer (DMC).

• System data acquisition concentrator (SDAC).

• Engine interface unit (EIU). The ECAM system receives engine and aircraft data and displays this on two cathode ray tubes (CRTs). The ECAM system is designed to give the flight crew primary and secondary engine/aircraft data. The flight warning system monitors all data that relates to a class ONE indication. This is regarded as the highest priority type annunciation. The system display can be transferred to the navigation display (ND) CRT by a selector switch.


© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications


© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications ECAM Indications The ECAM system displays both engine and aircraft data. The upper and lower ECAM CRTs display engine/aircraft data in digital and analogue form. Upper ECAM CRT The upper ECAM CRT will display the following engine and aircraft data:

• EPR command.

• EPR actual.

• EGT.

• N1 rotor shaft speed.

• N2 rotor shaft speed.

• Fuel flow.

• Fuel on board (FOB).

• Slat and flap position. The upper ECAM CRT display is also used to give warning information of class ONE alerts. This is given in the form of a message. Note: A1 series of engines have bump switches to enhance the take off performance. Whenever the bump is selected the alpha B will appear next to the EPR gauge. The B will disappear when:

• Mn of 0.45 is reached.

• The aircraft altitude has reached 15000 ft. Lower ECAM CRT The lower ECAM CRT will display the following engine and aircraft data:

• Fuel used.

• Oil quantity.

• Oil pressure.

• Oil temperature.

• Engine vibration for N1 and N2.

• Nacelle air temperature (NAC).

• Total air temperature (TAT).

• Static air temperature (SAT).

• Aircraft gross weight. Note: The NAC will appear on the lower ECAM CRT depending on modification standard of the aircraft. Pre mod standard indicates NAC only when an exceedance has occurred of 320°C. Post mod standard has the NAC indicated all the time. During engine start up the start air valve position, bleed air pressure and igniter selection are displayed in the NAC position.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications ECAM Indications Upper CRT Engine pressure ratio (EPR) Actual EPR indication is green. EPR limit is the thick amber index. EPR TLA angle is the white circle. Transient EPR is the blue arc. Idle indication flashing green for 10 seconds then steadies for both engines at idle in flight. REV indication for thrust reverser status. Exhaust gas temperature (EGT) Actual EGT indication is normally green. When EGT exceeds 610 deg.c the indication remains green the pointer pulses amber. The values pulse red when EGT at red line. EGT over-temperature is the red mark. If an over-temperature occurs a red mark appears at the max value achieved. It will disappear after a maintenance action through the MCDU. Max permissible EGT red line at beginning of red arc. During engine start the max permissible will be at starting value. Max EGT is the thick amber index. This is not displayed during engine start.

N1 Actual N1 indication is normally green. Pulses red when N1 limit is exceeded. Pulses amber when N1 exceeds N1 rating limit in N1 mode. Max permissible N1 is the red line indication at beginning of the red arc. N1 overspeed occurs a red mark appears at the max value achieved. It will disappear after a maintenance action through the MCDU. N2 Actual N2 indication is normally green. N2 goes red when limit is exceeded also a red cross appears next to the digital value. It will disappear when after a maintenance action through the MCDU. N2 indication is highlighted and boxed grey during engine start sequence. Thrust limit mode TOGA, FLX, MCT, CL and MREV are displayed in blue. EPR rating limit is displayed in green. De rate temp indication is displayed in blue. Actual fuel flow Actual fuel flow is displayed in green and gives real time indication of fuel flow for left and right engines.




DET

V250

377

UPPER ECAM CRT DISPLAY

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications ECAM Indications Lower CRT Fuel Used Fuel used indication is normally green. Freezes at last value when engine is shut down and resets at next engine start. Last two digits are dashed if fuel used indication is inaccurate due to loss of fuel flow for 1 minute. Oil quantity Oil quantity indication is normally green. At 5 quarts the advisory level is reached and the indication pulses. At 7 quarts and above the pulsing stops. Oil pressure The oil pressure indication is normally green. The indication pulses if the oil pressure exceeds 390 psi increasing or 385 psi decreasing. Between 80 and 60 psi the indication is amber. Below 60 psi the indication is red. Oil temperature Oil temperature indication is normally green. The indication pulses above 156 deg.c increasing and 150 deg.c decreasing. The indication becomes amber with an ECAM warning if temperature exceeds 165 or above 156 for more than 15 minutes or the temperature is below minus10 deg.c.

Ignition and start valve position The ignition and start valve positions are displayed during start up only. The selected igniters are displayed in green. The bleed pressure indication is normally green. If the pressure goes below 21 psi or suffers an over pressure the indication is amber as long as the start valve is not closed. Nacelle temperature (NAC) The pre mod standard aircraft NAC is not normally displayed. It will appear pulsing green (advisory) when an exceedance has occurred. The post mod standard aircraft the NAC indication is there all the time displayed in green Vibration Vibration indication is normally green. The indication pulses green if vibration is above 5.0 units (advisory). Oil filter and fuel filter No indication if both filters are normal. The message CLOG will appear in amber when the differential pressure across the filter has been exceeded. An ECAM message will also be generated.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications ECAM System Fault Monitoring Purpose The ECAM system is designed to constantly monitor for engine /aircraft parameter deviations from the normal. The deviation can then be annunciated to the flight crew. Description Normal parameter indication is:

• Green. Approaching parameter deviation the indication is:

• Flashing green. Warning condition (Class 1 Level 3) The parameter deviation indication is:

• Steady red indication.

• Master warning light on glare shield.

• Repetitive audible chime.

• ECAM message. Caution condition (Class 1 Level 2) The parameter deviation indication is:

• Steady amber indication.

• Master caution light on glare shield.

• Audible chime.

• ECAM message.

ECAM Messages The ECAM displayed messages are enunciated to the flight crew in the order of priority. The alert level classification for faults is as follows:

• Class 1 Level 3 Red warning with repetitive chime.

• Class 1 Level 2 Amber caution with chime.

• Class 1 Level 1 Amber caution with no chime. The upper ECAM CRT will display all warning type messages that are generated. This will display in the left memo box. The lower ECAM CRT will display messages of caution and status. The Lower ECAM CRT also has the facility to display other systems of the engine and aircraft.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications ECAM System Pages Purpose The ECAM pages give the flight crew a detailed parameter and system status of the aircraft and engine systems. These pages can also assist in troubleshooting. Description There are twelve pages of information covering the systems of the engine and aircraft. The pages are as follows:

• Engine.

• Bleed air.

• Cabin pressure.

• Electrical.

• Hydraulics.

• Fuel.

• Auxiliary power unit (APU).

• Conditioning.

• Doors.

• Wheels.

• Flight controls.

• Engine/air. The pages can be called up by either the flight crew manually or automatically according to the flight phase the aircraft is in.

ECAM Status Page The status page has information of faults that affect the redundancy of a system. Status messages do not directly affect the aircraft operation but reference to the MMEL is required before aircraft despatch. By depressing the status select button the status screen will appear. Status information that has occurred during flight will be alerted to the flight crew by a pulsing STS on the upper ECAM CRT warning memo box. This occurs when the engines are shut down.




DETV250381

ECAM AIRCRAFT/ENGINE SYSTEM PAGES

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications ECAM Flight Phase Displays During certain conditions of aircraft operation the lower ECAM CRT is configured to display certain pages of information. The page selection is automatic and dependant on what flight phase the aircraft is operating in. The flight phase selection is done automatically. If at any time a different page of information is required by the flight crew this can be selected via the systems page manual select panel. The flight phases are numbered 1 through to 10. Each flight phase requires a certain page of information to be displayed. The flight phases are as follows: Flight phase 1 Aircraft electrical power up. Systems page DOOR/OXY. Flight phase 2 Engine start to minimum idle. The ENGINE page will be displayed during engine start. The WHEEL page will be displayed after 2nd engine start. The FLT/CTL page replaces the wheel page for 20 seconds if the side sticks are moved or the rudder is deflected by more than 22 degs. Flight phase 3 Engines to power level above idle. Systems page ENGINE.

Flight phase 4 Aircraft speed in excess of 80 kts. Systems page ENGINE. Flight phase 5 Aircraft lift off. Systems page ENGINE. Flight phase 6 Aircraft above 1500ft. Systems page CRUISE. The cruise page appears when the slats are in and the engines are no longer at take off power. The cruise page disappears when the landing gear is selected down. Flight phase 7 Landing gear down. The aircraft is below 600 ft. Systems page WHEEL. Flight phase 8 Aircraft touch down. Ground spoilers are displayed at extended position. Systems page WHEEL. Flight phase 9 Aircraft below 80 kts. Landing inhibit message disappears. Systems page WHEEL. Flight phase 10 Aircraft at the gate. Both engines shut down. 5 minutes after 2nd engine shut down the FWC starts a new flight leg in phase 1.




ECAM FLIGHT PHASE DISPLAYS DET

V250

380

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications Flight Deck Centre Pedestal The following are some of the Engine related controls and interfaces: 1. Captains Multipurpose Centralised Display Unit

(MCDU) 2. Systems Display Control Panel 3. First Officers Multipurpose Centralised Display Unit

(MCDU) 4. Engine No. 2 Thrust Lever 5. Engine No. 2 Master Switch 6. Ignition Mode Selector Switch 7. Printer 8. Engine No. 1 Thrust Lever 9. Engine No. 1Master Switch




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications Flight Deck Overhead Panel The following are some of the Engine related interfaces found on the Overhead Panel: 1. N1 Mode Selector Switches for No. 1 & No. 2 engine 2. Engine Manual Start Switches for No. 1 & No. 2 engine 3. Engine and Auxiliary Power Unit (APU) Fire Panel 4. FADEC ground power switches for No. 1 & No. 2 engine




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications Shaft Speed Indicating System Purpose The speed indicating system provides signals of:

• N1 shaft speed.

• N2 shaft speed. The indications are used for:

• The ECAM CRT display.

• EEC control. A dedicated signal is used for trim balancing purposes. Location The N1 speed sensors are located in the front bearing chamber mounted on the no.2 bearing support. The N2 speed indication is the output signals from the dedicated generator. Description The speed indicating description is as follows: N1 System Three pulse probes supply the N1 indication. The pulse probes operate by monitoring the passage of a phonic wheel. The phonic wheel passage across the pulse probe generates an output signal relative to a percentage of a revolution.

Hence if the phonic wheel has 60 teeth then 60 pulses represents a complete revolution of the N1 shaft. N2 System The N2 indication is supplied by a dual output signal from channel B of the dedicated generator.

• An output goes to the channel B side of the EEC.

• An output goes to the EVMU. Fan Trim Balance The fan trim balance probe is located in the same place as the speed pulse probes. This probe supplies a dedicated signal for monitoring of LP system unbalance. The probe is also different from the speed probes. It cannot be utilised to give N1 speed indication. The pulse probe monitors a datum tooth of the phonic wheel. This tooth is in line with the no.1 fan blade.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications N1Speed Indicating System Operation The probes comprise of two pole pieces, a permanent magnet, and a coil wound on to one of the pole pieces. The pole pieces span two teeth of the phonic wheel. The phonic wheel is an integral part of the fan stub-shaft and has 60 teeth. As the shaft rotates and the teeth of the phonic wheel pass the pole pieces and a voltage pulse is produced in the winding. The number of pulses produced is directly proportional to the speed of the shaft. This signal is passed to the EEC and is used to display N1 speed on the flight deck and also for the engine control circuits as required. Trim Balance Probe The signal from this probe is only used during trim balance operations and provides the phase relationship between any out of balance forces present and a datum position. The trim balance probe senses the passage of one specially modified tooth on the phonic wheel and produces one pulse per revolution.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications Changeover of Speed Probe Harnesses AMM ref. 77-11-00-860-010 The following procedure outlines the requirement for connecting to the spare N1 probe terminals when a signal failure of N1 has occurred. Changeover procedure • Carry out the flight deck checks as per aircraft

preparation as advised by the AMM.

• Open the fan cowl doors (71-13-00-010-010).

• Deactivate the thrust reverser HCU (78-30-00-040-012).

• Open the thrust reverser C ducts (78-32-00-010-010).

• Remove the hose from the upper ignition unit. This will allow access to be gained to the terminal connections.

The terminal connectors are numbered and are in pairs. The pairing is as follows: Channel A speed probe no.1 is connected to terminals no. 1 and 2. Channel B speed probe no.3 is connected to terminals no. 5 and 6. Speed probe no. 2 is at the spare terminals that are no. 3 and 4. The trim balance probe is connected to terminals no. 7 and 8.





© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications Exhaust Gas Temperature (EGT) Indicating System Purpose EGT is displayed to the flight deck via the ECAM system to give the flight crew an indication of the engine temperature. This allows the engines to be operated within the temperature limitations as advised by IAE. Location The EGT thermocouples are located at the exhaust outlet. The EGT T/C leads come together at a junction box located at BDC of the turbine casing. Description The EGT is measured by 4 thermocouples, which are located in the support struts of the turbine exhaust case (engine station 4.9). The 4 thermocouples are connected to the junction box by a thermocouple harness. The materials used for the thermocouples and harnesses are:

• Chromel (CR).

• Alumel (AL). An extension harness connects the EGT junction box to channels A and B of the EEC.

Indication The EGT indication appears on the upper ECAM display unit. The ECAM provides the EGT indication: • In analogue dial gauge format. • In digital format. EGT is below 610 deg.c. The actual EGT indication is normally green. EGT is > 610 deg c. The indication pulses and changes colour to amber. EGT is > 635 deg c. • The indication becomes red. • The MASTER WARN light comes on, accompanied by

the repetitive audible chime. The following message appears on the ECAM upper CRT: EGT OVERLIMIT • The maximum value reached is memorised. • A small red line remains positioned on the analogue

scale at that value (max pointer). Note: The small and large nuts that secure the EGT leads to the junction box must torque check and tightened during the A check until further notice.(ref SB 77-0009) Single and dual channel failures have occurred due to loose EGT securing nuts.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications P3/T3 Sensor Purpose To give the EEC an input signal of;

• P3 pressure for fuel scheduling and surge detection.

• T3 temperature for trend monitoring. Location The P3/T3 sensor is located on the combustor casing at the one o’clock position. Description The P3/T3 sensor is a dual-purpose aerodynamically shaped probe. It measures the pressure and temperature of the air stream at the inlet of the diffuser case. The resultant data is transmitted to the EEC for control purposes. At the EEC the pressure enters a transducer. The temperature signal is received as a resistance value.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications Engine Pressure Ratio (EPR) Indicating System Purpose To indicate to the flight deck a parameter that is the representation of engine thrust. Location The main components of the EPR system are the P2/T2 probe and the P 4.9 pressure rakes. They are located: P2/T2 probe in the intake cowl at approximately TDC. P4.9 pressure rakes are in the exhaust duct of the LPT. Description The engine pressure ratio (EPR) is used to set and control the engine thrust EPR. EPR is:

P4.9 P2

The P2/T2 Probe measures P2. A pressure rake measures P4.9. The pressures from these sensors are routed to the EEC. The EEC processes the pressure signals to form actual EPR and transmits the EPR value to the ECAM for display on the upper screen. Each of the two EEC channels carries out this operation independently.

EPR Indications The actual EPR is displayed in green. The associated indications are:

• EPR maximum has a thick amber line.

• The maximum EPR value corresponds to thrust limit mode, which can be any one of the five conditions that follow;

Take off/go around mode (TO/GA). Flexible take-off mode (FLX). Maximum continuous thrust mode (MCT). Climb mode (CLB).

Flex TO Temperature is an assumed temperature entered by the flight crew through the MCDU to the FMS facility.

EPR reference is the predicted EPR value according to TRA.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications EPR System P2/T2 Sensor Purpose The P2/T2 sensor is a dual-purpose probe, which measures the total air temperature and pressure in the inlet air stream. The temperature and pressure signals are fed to the EEC. Location The sensor is installed at the 11 o'clock position in the air inlet cowl. Description The temperature is measured by two platinum resistance elements. Each channel of the EEC monitors one of the elements. The pressure signal is fed to a pressure transducer in the EEC. The sensor is electrically heated to provide anti ice protection. The EEC software corrects any temperature signal errors caused by heating. Note: The probe anti icing heater utilises 115V AC from the aircraft electrical system. Higher thrust engines (V2533) have a longer probe. The relay box on the right hand side of the fan case controls the selection of voltage to the probe heater unit.

EPR System P4.9 Rake Purpose The P4.9 rakes send a pressure signal to the EEC for the EPR system. Location The P4.9 rakes are located in the exhaust OGV’s. They are in the 3, 6 and 9 o’clock position. Description The P4.9 pressure rakes send a pressure signal down a common tube to a transducer within the EEC.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications Engine Vibration Indicating System Purpose The system monitors engine vibration for engine 1 and engine 2. Location The vibration transducer is located on the engine fan case in the 11 o’clock position. Description A vibration transducer on each engine fan case does monitoring. This produces an electrical signal in proportion to the vibration detected and sends it to the engine vibration-monitoring unit (EVMU). Two channels come from each engine. The EVMU provides signals of:

• Vibration.

• N1 (LP shaft speed).

• N2 (HP shaft speed). These are displayed on the engine page of the ECAM. The vibration transducer is installed on the fan case at the top left side of the engine. It is attached with bolts and is installed on a mounting plate.

Indications The engine vibration indications are displayed in green on the lower ECAM display unit on the engine and cruise pages. The ECAM display unit receives the information through the ARINC 429 data bus via the SDAC 1 and SDAC 2. If the advisory level is reached, the indication flashes (0.6-sec bright, 0.3-sec normal). If the indication is not available, 2 amber crosses replace the corresponding indication. Note: A5 engines have a dual cable. D5 engines have a single cable.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications P2 and T2 Probe Harness

.

The P2/T2 harness electrical connections are shown below




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance l Engine Indications Temperature Measurement Harness

The temperature measurement harness electrical connections are shown below.




SECTION 14

ENGINE STARTING AND IGNITION (Chapter 80)

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System

Page 14-1 Revision 2 • Automatically.

Engine Starting and Ignition System Purpose The engine starting system provides the power which turns the HP rotor to a speed at which an engine start can occur. The engine ignition system provides the electrical spark that is required to ignite the fuel air mix in the combustor. The ignition system is used for:

• Engine starting on ground and in flight.

• Prevention of a flame out by providing a continuous spark during engine running when the aircraft is flying through a heavy rainstorm for example.

Description The system comprises of the following:

• Pneumatic starter motor.

• Starter air control valve.

• Dual ignition system.

• Pneumatic ducting.

• Start control panels on the flight deck for auto starts, manual starts and starter motor operation.

• ECAM indications. Starting of the engine for the Airbus A319, 320 and 321 can be done either:

• Manually.

Basic start sequence Whichever method of starting is used the control is either from the EEC or from the cockpit through the EEC. In both cases the start sequence is initiated in the flight deck. Upon selection for engine start an electrical signal is sent to open the starting air valve. The starting air valve opens and admits an air supply into the starter motor. The starter motor rotates the high speed external gearbox, that in turn rotates the radial drive shaft (tower shaft), that in turn rotates the HP system (N2). As the HP system spools up the LP system starts to rotate due to the induced airflow.

• At 10% N2 the dedicated generator comes on line.

• The HP system is rotated for 30 seconds to remove rotor bow.

• After 30 seconds the fuel and ignition are selected on.

• When ignition of the fuel takes place the engine accelerates towards minimum idle

At 43% N2 the starter air valve is deselected, by the EEC. At 50% N2 and above the EEC auto start protection is cancelled. The engine idles at approximately 60% N2. Note: Above 50% N2 the command for engine shut down is done from the master lever only.



© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Starter Air Duct Purpose To provide a means of supplying air to the starter motor. Location The starter air duct is located on the right hand side of the engine fan casing (intermediate module). Description Air supplies for the pneumatic starter motor may be supplied from:

• The aircraft APU.

• Cross bleed from the other engine if already running.

• Ground starter trolley. Minimum duct pressure for starting should be between 30 and 40 psi. All ducting in the system is designed for high pressure and high temperature operation. Gimbal joints are incorporated to permit working movement. E-type seals located between all mating flanges prevent air leakage; Vee-band coupling clamps secure mating flanges.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Starter Air Control Valve Purpose The starter air control valve is designed to control the admittance of air to the starter motor. The valve is commanded from the flight deck via the EEC. Location The starter air control valve is located on the right hand side of the engine fan casing (intermediate module). Description The starter air control valve consists of the following:

• Butterfly valve for airflow control.

• Pneumatically operated.

• Microswitch position indication for valve positional status.

• Air filter to prevent valve operating mechanisms from contamination.

• Failsafe position of the valve is closed.

• Provision of a manual override for abnormal start attempts.

The starter air control valve is a pneumatically operated, electrically controlled shut-off valve. The valve is positioned on the lower right hand side of the engine fan casing.

The starter air valve controls the airflow from the air ducting to the starter motor. The start valve basically comprises a butterfly type valve housed in a cylindrical valve body with in line flanged end connectors, an actuator, a solenoid valve and a pressure controller. Manual operation The starter air valve can be opened/closed manually using a 0.375 in square drive. Access is through a panel in the right hand side fan cowl door. A valve position indicator is provided on the valve body. A micro switch provides valve position feed back information to the EEC.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Starter Air Valve Operation The starter air valve operation for opening and closing is as follows: Engine shutdown With no pneumatic air supply available the valve is spring loaded to the closed position. Valve opening Air upstream of the butterfly valve is filtered and routed through an orifice in the solenoid valve. Air upstream of the orifice is also admitted to the smaller piston of the double acting actuator. When the solenoid is energised the ball valve opens to admit air to the larger piston whilst simultaneously closing the vent port. The air acting on the larger piston overcomes the combined force of upstream air pressure acting on the smaller piston and the actuator spring. Movement of the actuator is translated through the linkage to rotate the butterfly valve towards the open position.

Valve closing When the solenoid is de-energised, at approximately 6000 rpm (43%) N2, the ball valve closes and air acting on the larger piston are vented to atmosphere through the vent. Air pressure and actuator spring pressure acting on the smaller piston then closes the butterfly valve. Any loss of air pressure will cause the butterfly valve to close under the action of the actuator spring.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Pneumatic Starter Motor Purpose The purpose of the pneumatic starter motor is to provide an initial rotational input to the high speed external gearbox in order to assist the engine to achieve a stable minimum power condition (low idle). Location The starter motor is located on the front face of the high speed external gearbox. Description The starter motor consists of the following:

• Oil filler/level plug.

• Drain plug with a built in magnetic chip detector.

• QAD devices to allow for ease of maintenance. The starter motor gears and bearings are lubricated by an integral lubrication system. A quick attach/disconnect adapter (QAD) attaches the starter motor to the external gearbox. A quick detach Vee clamp connects the starter motor to the adapter. Note: There are two standards of starter motor available for the V2500 Powerplant. The current being the synchronous clutch engagement unit. The synchronous clutch allows for smoother crash engagements thus reducing the wear and damage caused by such operations.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Starter Motor Operation The starter is a pneumatically driven turbine unit that accelerates the HP rotor to the required speed for engine starting. The starter comprises of the following:

• A single stage turbine.

• A reduction gear train.

• A clutch and an output drive shaft. These are all housed within a case incorporating an air inlet and exhaust. Compressed air enters the starter, impinges on the turbine blades to rotate the turbine, and leaves through the air exhaust. The reduction gear train converts the high speed, low torque rotation of the turbine to low speed, high torque rotation of the gear train hub. The ratchet teeth of the gear hub engage the pawls of the output drive shaft to transmit drive to the external gearbox, which in turn accelerates the engine HP compressor rotor assembly. When the air supply to the starter is cut off, the pawls overrun the gear train hub ratchet teeth allowing the turbine to coast to a stop. The engine HP turbine compressor assembly, the external gearbox and starter output drive shaft continue to rotate.

When the starter output drive shaft rotational speed increases above a predetermined rpm, Centrifugal force overcomes the tension of the clutch leaf springs, allowing the pawls to be pulled clear of the gear hub ratchet teeth to disengage the output drive shaft from the turbine.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Engine Ignition System Purpose The ignition system is designed to provide the means of igniting the air/fuel mix in the combustor. Location The ignition system units are located in the following positions:

• The relay box is located on the right hand side of the engine fan case.

• The high energy ignition units (HEIU's) are located on the right hand side of the core engine. Mounted on the HPC casing.

• The igniter plugs are located on the combustion diffuser casing at fuel spray nozzle positions no. 7 and 8.

Description Two independent ignition systems are provided. The system is made up of the following units:

• Ignition relay box.

• Two ignition exciter units.

• Two igniter plugs.

• Two air cooled HT ignition connector leads.

The ignition system can operate in various modes. These modes are as follows: Dual igniter select:

• All in flight starts.

• Manual start attempts.

• Continuous ignition. Single alternate igniter select:

• Auto starts. Continuous ignition select:

• Engine anti ice.

• Take off.

• Approach.

• Landing.

• EIU failure. Continuous ignition may also be selected manually. The ignition exciters provide approximately 22.26 Kv and the igniter discharge rate is 1.5/2.5 sparks per second. Test Operation of the ignition system can be checked on the ground, with the engine shut down, through the maintenance menu mode of the CFDS.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Ignition Relay Box Purpose Used for connection and the isolation of the high energy ignition units. Location The relay box is located on the right hand side of the engine fan casing. Description The ignition system utilises 115V AC supplied from the AC 115V normal and standby bus bars to the relay box. The 115V relays, which are used to connect/isolate the supplies are located in the relay box and are controlled by signals from the EEC. Note: The same relay box also houses the relay that controls the 115V AC supplies for P2/T2 probe heating.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Engine Starting Electrical Control Engine Interface Unit (EIU) Purpose The EIU is an interface concentrator between the aircraft and FADEC system. Location The EIU is located in the avionics bay. Description There are two engine interface units (EIU's), one for each engine. The EIU is an interface concentrator between the aircraft and FADEC system. The EIU main functions are:

• To concentrate data from the flight deck panels.

• To ensure the segregation of the two engines.

• To provide the EEC with electrical power supply.

• To give the necessary logic and information between the engine and the aircraft systems.

• Receives discrete electrical signals from the cockpit. Digitises these signals and transmits them to the EEC. Also sends discrete signals to close air conditioning pack flow valves and increase the airflow from the APU if required.

Electronic Engine Control (EEC) Purpose To provide electronic signals for FADEC system unit control. Location The EEC is located on the engine fan casing right hand side. Description The EEC is the heart of the FADEC system and has control of the FADEC system components and constantly monitors their performance. The EEC will make adjustments where necessary to optimise the operation of the engine. During the starting of the engine the EEC generates the pneumatic starter valve opening/closing signal in respect of control switch selection (rotary selector, master lever, ‘MAN START’ push button switch) and N2 speed signal. The EEC will send any warning or caution message to the flight warning computer (FWC). The FWC will send this to the display management computer (DMC) for indication on the ECAM upper or lower CRT.





Page 14-19 Revision 2 • Be aware of the hazard radius area about the air

intake. Engine Ground Operation Safety Zones During run up operations, extreme care should be exercised when operating the engines. Purpose The purpose of ground operation of an aero engine is to prove the integrity for continued use. Diagnosis of faults can be determined through ground operation. General procedures • Apply the brakes and position the wheel chocks.

• Inspect the ground run area for loose debris.

• Avoid obstructing the air intake area.

• Head the aircraft into the wind wherever possible. The AMM will advise if this is not always possible.

• Cross wind conditions may cause parameter fluctuations in adverse conditions.

• Cross wind conditions can cause the engine to surge. A roaring type noise is evidence of an unstable condition that can lead to surge.

• Be aware of the jet wake generated with the engine running.

• There are minimum and maximum safe distances for power conditions between low idle and take off. The AMM will advise.

• Be aware of the noise hazard. Jet noise can seriously damage the hearing. The AMM advises that the appropriate hearing protection be worn.

• It is advisable to carry out an engine passages inspection prior to engine running.

• Ground running in icing conditions requires the use of the anti icing system. Icing conditions exist when the OAT is 5.5 deg.c or less with visible moisture present.

• Engine running should be kept to a minimum. The AMM advises for engine warm up, operation at high power, throttle movement rates and engine cool down prior to shut down.

• Be aware of the imbalance caused by single engine high power running. The AMM advises on the conditions of engine running required of the opposite engine.

Refer to the diagram below, which illustrates the inlet suction hazard areas for the conditions at idle and take-off thrust.



© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Safety Zones - Jet Wake Hazard AreasDuring run up operations, extreme care should be exercised when operating the engines. Refer to the diagram below, which illustrates the jet wake hazard areas for the conditions at idle and take-off thrust. Noise Danger Areas All persons working near the engine while it operates must wear ear protection. Loud noise from the engine can cause temporary or permanent damage to the ears




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Engine Operating Limits

Engine Rating N1 N2 EGT Max EGT Cont. EGT Start Pre start EGT

N1 Vib N2 Vib

V2533-A5 5650 14950 650 610 635 250 5.0 5.0

V2530-A5 5650 14950 650 610 635 250 5.0 5.0

V2528-D5 5650 14950 635 610 635 250 5.0 5.0

V2527-A5 5650 14950 635(E/M) 610 635 250 5.0 5.0

V2525-D5 5650 14950 620 610 635 250 5.0 5.0

V2500-A1 5465 14915 635 610 635 250 5.0 5.0

V2524-A5 5650 14950 635 610 635 250 5.0 5.0

V2522-A5 5650 14950 635 610 635 250 5.0 5.0E is enhanced performance. M is for the corporate A319 jet. The following operating limits apply to all engine ratings for the oil system.

Min start Min to 1.3EPR

Min to T/O Max trans Max limit Minimum Maximum

Oil Pressure 60 psi ISAdependant

Oil temperature -40 deg.c -10 deg.c 50 deg.c 156 deg.c amber

165 deg.c red


ENGINE GROUND OPERATION




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Engine Operation Purpose The purpose of engine operation on the ground is to validate non FADEC detected faults (mechanical failures) and to prove the integrity of an LRU or system after maintenance has been carried out. Flight deck The flight deck is where the engines are operated. There are two panels utilised for starting the engine. These panels are:

• The auto mode select panel (auto starts).

• The manual start panel. Auto mode select panel The auto mode select panel has the following:

• Two main engine master switches.

• Rotary switch. The engine master switches have two positions:

• Cut off.

• On. These switches activate the start air valve and the FMU, via the EIU and EEC, when in the auto mode. These switches activate the FMU, via the EIU and EEC, when in the alternate (manual) mode.

The rotary switch has three positions:

• Crank.

• Mode norm.

• Ign/start. The crank position allows operation of the starter motor only. The mode norm sets the ignition to auto function, for example when anti icing is selected the ignition comes on to a continuous operation. The ign/start allows the engine to be started. This switch must be in the ign/start position before selecting the master levers to the on position. Manual start panel. The manual start select panel allows the engine to be started in the non-auto function or manual mode. Note: The EEC software has a fuel flow reduction capability upon the detection of a stall. This is known as fuel depulse. The de-pulse logic is designed to assist the engine in recovery from a stall during starting.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Engine Operation Dry Motoring The starter motor may be operated under the following conditions only:

• 2 off 2 minute cycles followed by a 1 off 1 minute cycle.

• 1 off 4 minute cycle. In all cases the engine N2 indication must be allowed to decay to zero before commencing the next cycle. A 30 minute pause must be allowed for cooling before recommencing either method of the duty cycles. The V2500 engine can be dry motor operated by carrying out the following: Pre start checks • Thrust levers at idle.

• Master switch set to off.

• Auto mode selector set to normal.

• Manual start push buttons set to off.

• Aircraft booster pumps set as necessary. Dry motor procedure • Select the crank position on the auto mode select

panel.

• Select manual start push button on.

• Observe the engine rotor speeds for correct indication.

• Check engine vibration for within limits for engine start.

Engine Operation Wet Motoring The V2500 engine can be wet motor operated by carrying out the following: Pre start checks • Thrust levers at idle.




• Aircraft booster pumps set as necessary. Wet motor procedure • Select the crank position on the auto mode select

panel.

• Select manual start push button on.


• Check engine vibration for within limits for engine start.

• Select the engine master switch to the on position.

• Observe fuel flow indication. Note: In auto the EEC will deselect the starter cycle if the starter motor operation exceeds the starter cycle limit. In manual the de-selection of the starter motor must be done manually. If time has expired an ECAM upper caution message is displayed.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Engine Operation Auto Start AMM Ref. Ch 71-00-00-710-043 The V2500 engine can be started automatically by carrying out the following: Pre start checks • Thrust levers at idle.




• Aircraft booster pumps set to on. Auto start procedure • Select the ign/start position on the auto mode select

panel. By selecting the ign/start position the lower ECAM screen goes to the engine page.

• Select the engine master switch to the on position.


• Check engine vibration for within limits during the engine start.

• The HP system is rotated for 30 seconds to remove rotor bow.

• After 30 seconds the fuel and ignition are selected on.

• When ignition of the fuel takes place the engine accelerates towards minimum idle

The engine will not light up if the residual EGT is in excess of 250 deg.c. The EEC will continue the dry motor cycle until the temperature falls below 250 deg.c. At 43% N2 the EEC signals the starter motor to cut off.

• Above 50% N2 the EEC no longer has capability of closing the FMU. This function becomes sole priority of the flight crew.

• At idling conditions check that the indicated parameters are within acceptable limits.

If at any time the engine experiences a non normal event such as:

• Hot start.

• Stall.

• No N1 or N2 indications.

• Starter valve failure.

• Ignition failure.

• PRSOV failure. The EEC will abort the start sequence. Manual start AMM Ref. Ch 71-00-00-710-047 Using the manual start push buttons to do a manual start. The procedure requires that the manual start push buttons should be selected on before selecting the master switch to the on position. During the manual start the EEC does not have auto shut down priority all non normal events have to be monitored by maintenance personnel.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Failure to Start ECAM Indications Purpose The failure indications alert the flight crew as to the problem in hand and what to do as a reaction. Location The alert messages are displayed on the upper ECAM CRT. Description The following messages can be experienced on the upper ECAM CRT if a fault occurs:

• Fuel PRSOV not open in auto mode.

• Fuel PRSOV not open in manual mode.

• Starter time exceeded in auto mode.

• Starter time exceeded in manual mode.

• Start valve not open fault.

• Start valve not closed fault.

• Ignition fault in automatic mode.

• Ignition fault in manual mode.

• EGT overlimit and stall fault in automatic mode.

• EGT overlimit and stall fault in manual mode. All fault messages will generate a caution message to ECAM, an aural tone will be heard and the master caution light will be illuminated on the glareshield panel.

Each message that is generated both in auto and in manual will also be accompanied by further messages advising the flight crew on the actions required as a result of the failure. For example; ENG 2 FUEL VALVE FAULT -FUEL VALVE NOT OPEN -IF NO ENG LIGHT UP: -ENG MASTER 2------------------OFF This typical message that can be generated to the ECAM upper CRT. The message is for an auto start problem. ENG 2 FUEL VALVE FAULT -FUEL VALVE NOT OPEN -IF NO ENG LIGHT UP: -MAN START------------------------OFF -ENG MASTER 2-------------------OFF This is a typical message that can be generated to the ECAM upper CRT. The message is for a manual start problem.





71-00-00-700-013 Normal engine automatic start procedure71-00-00-700-015 Test No 1 dry motor leak check 71-00-00-700-016 Test No 2 wet motor leak check 71-00-00-700-017 Test No 3 Idle leak check 71-00-00-700-018 Test No 4 oil system static leak check 71-00-00-700-020 Test No 6 EEC system idle check 71-00-00-700-021 Test No 7 reserved 71-00-00-700-022 Test No 8 vibration survey 71-00-00-700-037 Normal engine manual start procedure 71-00-00-700-039 Test No 9 LP compressor (fan) trim

balancing – one shot method 71-00-00-700-038 Test No 9A LP compressor (fan) trim

balancing – trial weight method 71-00-00-700-024 Test No 10 performance test 71-00-00-700-025 Test No 11 high power assurance test 71-00-00-700-027 Test No 13 pre tested engine

replacement test 71-00-00-700-028 Test No 14 untested engine replacement

test 73-22-00-700-010 Operational test of the FADEC system

on the ground 73-22-34-710-010 Operational test of the EEC 74-00-00-710-041 Operational test of the ignition system

with the CFDS

72-00-00-710 041 01

Operational test of the ignition system without the CFDS





© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Starter Air Valve HarnessThe starter air valve harness connections are shown below.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Engine Starting and Ignition System Ignition System HarnessThe ignition system harness electrical connections are shown below.




SECTION 15

THRUST REVERSER SYSTEM (Chapter 78)

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Thrust Reverser System Introduction Purpose The thrust reverser is designed to assist the aircraft in decelerating quickly and safely upon landing. It also assists deceleration during an aborted takeoff. Location The thrust reverser is an integral part of the C duct assembly. The C duct assembly is mounted to the aircraft strut by four hinged brackets located at the top of the C ducts. They are held in the closed position by six latch locks located at the bottom of the C ducts. Description The reverser is a translating sleeve type system. It directs the fan air rearwards for normal forward thrust or forwards for thrust reverse. When the thrust reverser system is in the stow position the fan air exhausts at the common nozzle. This produces forward thrust. When the thrust reverser is deployed four linear motion actuators cause the translating sleeves to move rearwards. This moves the blocker doors from an axial to a radial position in the C duct fan exhaust area. The blocker doors forces the fan air through the cascades in a forward direction. The cascades are exposed whenever the thrust reverser is deployed.

Reverse thrust is selected from the flight deck by the gated reverse thrust levers. The EEC has control over the operation of the thrust reverse system. All signals to and from the thrust reverser are through the EIU and EEC.

Thrust reverser system features Electronic control. Hydraulic actuation system. Positional information feedback. Actuator lock position sensors and feedback. Electronic safety locks. Automatic restow system. Manual deployment and stow capability for maintenance. Manual lockout to allow aircraft to be despatched with an inoperative thrust reverser.


© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System


DE

TV25

0262

V2500 THRUST REVERSER SYSTEM V2500 THRUST REVERSER SYSTEM

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Thrust Reverser Assembly When the translating sleeves are in the forward thrust position the path of the fan bypass air is in the normal forward thrust. Rearward movement of the translating sleeves unveils the cascade deflectors and moves the blocker flaps from an axial to a radial position. This blocks the fan stream airflow and forces the fan efflux through the cascade deflectors.

Methods of deployment The thrust reverser can be deployed in one of two methods;

• Using the engine/aircraft hydraulic system. Moving the thrust reverse select levers that are mounted on the main forward levers does selection.

• Manual input by two hand turning points for maintenance purposes.

Safety features The thrust reverse system operation is controlled by the engine electronic control (EEC). The following are EEC controlled functions of the thrust reverse system. The thrust reverse system incorporates a double lock safety system to protect against inadvertent deployment. They are;

• Landing gear control unit (LGCU).

• Lock sensors on the locking actuators.

• EIU inhibit relays for uncommanded in-flight

deployment.

• Shut off valve that is signalled to operate from the SEC.

• The thrust reverser also has a system that will return the engine thrust to idle should the thrust reverse system inadvertently deploy.

• Auto restow is a system that is designed to stow the thrust reverser when an uncommanded deployment is detected.

Indications ECAM indications for fault annunciation of the thrust reverser system status are done by use of proximity sensors, relay select status, hydraulic system pressures and LVDT feedback signals. The locking actuator sensors detect unlocked conditions and the LVDT detects transient and deployed conditions. The signals are relayed from the EEC to the EIU and then to the ECAM screens.






In the flight deck this unlocked condition is identified as an amber coloured REV caption on the EPR indicating gauge.

Operation Thrust reverse is selected from the flight deck by pulling up on the thrust reverse select levers. The select levers are mounted on the front side of the main thrust levers. The thrust levers have a gated feature that allows thrust operation by the throttles in one direction only. The EEC has control of the thrust reverse system operation for deploy and stow. The EIU inhibit relay controls the DCV power signal for the control solenoid from the EEC to the DCV. Deploy Pulling up on the thrust reverse lever in the flight deck will send a signal for thrust reverse select to the EEC. This will also put the main throttles in the reverse thrust quadrant. The EEC will look for the following conditions before thrust reverse will be allowed; • The EEC will check that the aircraft is on the ground by

checking the LGCU signal of the aircraft computers. • The EEC will check that the engine is running by means

of a N2 signal. • The EEC cannot deploy the thrust reverser until the EIU

inhibit relay is active. • SEC control signal for the shut off valve. The hydraulic isolation valve solenoid and the directional control valve solenoid will both be energised for a deploy condition. This will admit high pressure hydraulic fluid to the stow and deploy sides of the thrust reverse system. The lower locking actuators will unlock and the EEC will see a signal from the proximity sensor of reverser system unlocked.

When the translating sleeves have moved to 78% of the full deploy position the amber REV indication will change to a green REV indication. When green REV is indicated the full reverse thrust power is available to the flight crew. Stow To stow the thrust reverse system the flight crew will return the throttles to the idle detent position and select levers to the down position. This will put the throttles back to the forward thrust quadrant. The hydraulic isolation valve solenoid is energised and the directional control valve solenoid is de-energised for a stow condition. This leaves high pressure hydraulic fluid present on the stow side of the system. As the translating sleeves move from deploy back to stow the flight deck indication will change from green to amber on the REV indication. When the thrust reverser has reached the fully stowed position the amber REV will go and the EPR gauge will return to normal indication. This will indicate that the thrust reverser is fully stowed and locked.



© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Controls and Indications Thrust reverse is selected from the flight deck by use of latching selector levers that are mounted on the main throttle control levers.

Controls Pulling the levers upwards will initiate the sequence of events that will deploy the reverser system. The EEC, in conjunction with the EIU, controls the deployment and stowing of the reverser system. The levers move into the thrust reverse quadrant therefor while in this position throttle movement is only possible in the thrust reverse mode. Movement towards the maximum throttle stop for thrust reverse is possible but the engine will only accelerate when the EEC has feedback of the translating sleeve status. The translating sleeve must be beyond 78% of the fully deployed position. Stowing the thrust reverser requires the latching select levers to be pushed down and the main throttles will revert back to normal forward thrust. This will also stow the reverser system.

Indications The thrust reverser system indications appear on the ECAM CRTs. The EPR indication is used to display the status of the thrust reverser.

Normal indication Thrust reverser stowed and locked.

REV in colour amber Thrust reverser unlocked and in transit.

REV in colour green Thrust reverser deployed. Thrust reverser indications of non-normal conditions will be indicated to the ECAM screens in the form of a message. The following are the associated ECAM messages that appear to the ECAM screens;

• Reverse Unlocked.

• Reverser Fault.

• Rev pressurised.

• Rev Switch Fault. By entering the CFDS screens the faults can be interpreted to pinpoint the location.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Hydraulic System The hydraulic system provides the force required to move the translating sleeves for both deploy and stow conditions. The hydraulic system comprises of the following;

• Linear motion actuators.

• Flex shaft.

• Hydraulic control unit comprising of a HIV and DCV.

Linear Motion Actuators There are four linear motion actuators per C duct set. The two upper actuators are non-locking and incorporate LVDT’s for feedback to the EEC. The two lower actuators are locking they incorporate proximity sensors to give indication to the EEC of lock and unlock conditions.

Flex shaft The four linear motion actuators are kept in synchronisation movement by flexible shafts that have a high torsion resistance.

Hydraulic Control Unit The hydraulic control unit comprises of the following items;

• Hydraulic isolation valve (HIV).

• Directional control valve (DCV).

The HIV controls the presence of high pressure hydraulic fluid in the thrust reverser system. The energising of a control solenoid valve controls this valve. For stow and deploy conditions this valve must be energised. The DCV controls the flow direction of the high pressure hydraulic fluid once it is in the thrust reverser system.

• The DCV will direct the high-pressure hydraulic fluid to both the stow and deploy sides of the system for deploy conditions.

• The DCV will direct the high-pressure hydraulic fluid to the stow side of the system for stow conditions.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Hydraulic Control Unit Purpose The hydraulic control unit (HCU) is designed to control the safe passage of the high-pressure hydraulic fluid to the thrust reverser system.

Location The HCU is located between the top of the engine fan case and the aircraft strut. Access is gained by opening the left hand side fan cowl door.

Description The HCU is a self contained LRU designed to control the flow of high pressure hydraulic fluid. The EEC and EIU has control over the HCU control solenoids. When the EEC detects a demand for thrust reverse operation both EEC and EIU will signal the HCU control solenoids. The HCU has the following features;

• A hydraulic isolation valve (HIV) which controls the flow of high-pressure hydraulic fluid into the thrust reverser system. A control solenoid valve controls the HIV function.

• A directional control valve (DCV) which controls the direction of flow of the high-pressure hydraulic fluid to either the deploy or stow sides of the system. A control solenoid valve controls the DCV function.

• A pressure switch to feedback system pressure status to the EEC.

• A filter with a clog indicator that gives a visual indication of the filter condition when it becomes contaminated.

• A bleed valve.

• Provision for locking out the valve operation for maintenance and flight.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Hydraulic Control Unit Operation Deploy When the thrust reverser select levers are moved to the up position the EEC detects that the thrust reverser system is required. • The EEC has selection control over the HIV. • The EIU has selection control over the DCV. The HIV control solenoid will be energised. This moves a lockpin away from the pilot valve chamber orifice. High pressure hydraulic fluid enters the left hand side of the chamber and forces the pilot valve to move to the right. The pilot valve moves against a spring that exerts a pressure on the right hand side of the pilot valve. The pilot valve recess moves in line with the hydraulic fluid supply tube. This admits high pressure fluid into the thrust reverse system and initially to the stow side of the system. The pressure switch moves to the high pressure indicating position. The DCV control solenoid will be energised. This moves a lockpin away from the pilot valve chamber orifice. High pressure hydraulic fluid enters the left hand side of the chamber and forces the pilot valve to move to the right. The pilot valve moves against a spring that exerts a pressure on the right hand side of the pilot valve. The pilot valve recess moves in line with the hydraulic fluid supply tube. This admits high pressure fluid into the thrust reverse system deploy side of the system. There is now pressure present in both the stow and deploy sides of the system.

The restrictors in the deploy supply tube to the DCV delay the pressure build up to the deploy side so the pressure present on the stow side can push the locking actuators towards the stow. This releases the pressure acting on the tine locking mechanism. Stow To stow the thrust reverser the select levers are moved to the down position. • The EEC has selection control over the HIV. • The EIU has selection control over the DCV. The DCV control solenoid is de-energised. The pilot valve moves to the right due to spring pressure alone. This leaves high pressure hydraulic fluid present in the stow side of the system. When the thrust reverser system has fully stowed the EEC will sense this by a feedback signal coming from the unlock sensors. The EEC will then de-energise the HIV control solenoid. The control solenoid pilot valve will move to the right due to spring pressure alone and the high pressure hydraulic fluid is cut off from the thrust reverser system. HIV Deactivation The deactivating lever prevents the HIV pilot valve from moving. This prevents high pressure hydraulic fluid from entering the thrust reverser system.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Lower Locking Actuators Purpose The hydraulic actuators in general control the movement of the translating sleeves. The lower locking actuators incorporate a locking mechanism which gives;

• A feedback signal to the EEC of actuator locked or unlocked.

• A means of preventing the translating sleeves from uncommanded movement.

Location The lower locking actuators are located in the thrust reverser C duct units at the lower positions.

Description The actuators in general control the movement of the translating sleeves in a linear motion. The lower actuators on either thrust reverser C duct have a locking mechanism incorporated in the design. The locking mechanism adds to the safety of the system. The lower locking actuators incorporate the following features;

• Hydraulically operated linear motion actuators.

• A locking mechanism called a tine lock. This can only be unlocked when hydraulic pressure is present in the system.

• A linear motion actuator with pressure surfaces on either side of the pressure plate.

• An acme screw thread that rotates a worm gear.

• A worm wheel that rotates the flex shafts.

• A manual unlocking feature for maintenance purposes.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Lower Locking Actuators Operation Deploy The EEC will energise the HIV and DCV control solenoids which allows high pressure hydraulic fluid to be present in the thrust reverser system. This is on the stow and deploy sides. The restrictor in the pressure feed tube to the deploy side delays the deploy pressure build up enough to allow the stow pressure to initially push the actuator piston towards the stow direction. This releases the lock pressure on the tine locking mechanism. The unlock sleeve is then pushed towards the right of the tine lock. With the locking sleeve clear of the tine lock the tine lock flexible spring type fingers are free to flex. When the locking sleeve moves a lever assembly also moves. The lever assembly is linked to the external area of the actuator. The lever has a target attached to it.

• When the actuator is at stow the target is in line with the proximity sensor.

• When the actuator is at deploy the target is away from the proximity sensor.

The EEC detects these conditions. High pressure fluid being present in both sides of the system forces the actuator to move towards the deploy direction. This bias of movement exists because the surface area of the deploy side of the pressure plate is greater that the surface area of the stow side.

Stow To stow the thrust reverser the high pressure fluid present on the deploy side of the system must be reduced to tank pressure and ported back to the hydraulic reservoir of the aircraft. The EEC will de-energise the DCV control solenoid and this will leave high pressure hydraulic fluid present on the stow side of the system. The actuator will now move in the stow direction. The head end of the actuator engages into the tine lock. The locking sleeve will move into position to immobilise the tine lock by spring pressure. As the locking sleeve moves to the lock position the target on the unlock indicator moves in line with the proximity sensor. The EEC detects this and sees that the thrust reverser system is stowed.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Upper Non Locking Actuators Description The hydraulic actuators in general control the movement of the translating sleeves. The upper non locking actuators incorporate a linear variable displacement transducer (LVDT) that feeds back translating sleeve status of translation and deployment. The actuators operate in the same manner as the locking actuators.

• For deployment the high pressure hydraulic fluid is present on both sides of the system.

• For stowing the high pressure hydraulic fluid is present on the stow side of the system only.

The LVDT monitors the movement of the translating sleeves and feeds back the signals to the EEC. This is done to tell the EEC that the translating sleeves are in transit and when 78% of travel towards the deploy has been achieved.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Shut Off Valve Purpose To give additional safety to uncommanded deployments of the thrust reverser. This is known as the third lock of safety. The shut off valve has a filter assembly installed in line with the valve.

Location The shut off valve is located in the aircraft strut at the front. It is behind the HCU.

Description The shut off valve has the following features; Control solenoid. Two position valve assembly.

Operation The shut off valve operation for opening and closing relies upon the signals from the spoilers and elevators computer (SEC). Pulling up the thrust reverse select levers will signal the SEC to open the shut off valve. The SEC sends a signal to the shut off valve relay. The relay energises the solenoid and opens the shut off valve. The high pressure hydraulic fluid now flows towards the HCU. To close the shut off valve the selection of thrust reverse levers must be in the forward thrust position.

Manual Bypass Non Return valve Purpose For normal thrust reverser operation provides a one way directional flow for the hydraulic fluid. For maintenance purposes it allows the flow of hydraulic fluid easier when manual deployment and stow of the thrust reverser is required.

Location The manual bypass valve is located in the aircraft strut just behind the shut off valve.

Description The manual bypass non return valve allows the flow of hydraulic fluid in one direction only. There is a very hard spring loaded valve inside that is difficult to unseat when carrying out manual operations of the thrust reverser system. The bypass handle allows the valve to become unseated for maintenance operations only. Access to the bypass valve is through an access panel located on the left hand side of the aircraft strut.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Thrust Reverser Flex (Synchronisation) Shaft Purpose The flex shaft purpose is to maintain synchronous movement of the actuators. This prevents any one actuator from moving faster than the others.

Location The flex shaft is located within the deploy tube system.

Description The flex shaft system comprises of the following; T piece housing assembly. This allows the distribution of high pressure hydraulic fluid to both sides of the thrust reverser system. Two flexible tubes. This allows the crossover shaft to link the reverser halves together while allowing the C ducts to be opened. Two rigid tubes. These are found between the upper and lower actuators. They carry hydraulic fluid to the deploy side of the system. Three flexible shafts. These link all the actuators together.

Note: The two deploy tubes have a telescopic coupling at one end to permit simple removal and installation without disturbing the actuators.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Thrust Reverser Deflector Boxes (Cascades) Purpose The cascades are designed to direct the fan air to provide the reverse thrust for the engine.

Location The cascades are located between the inner and outer translating sleeve sleeves. They are mounted on the fixed section of the C ducts.

Description There are 16 cascades fitted to the thrust reverser system. The cascades are designed to direct the fan air forwards thus providing for the function of the thrust reverse system. They are designed to direct the fan air away from the ground thus reducing the risk of debris from being blown up and ingested into the engine. They are designed to direct the fan air away from the airframe thus not inducing any unnecessary stress upon the airframe itself.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Operation Thrust reverse is selected from the flight deck by pulling up on the thrust reverse select levers. • The EEC has main control of the thrust reverse system

operation for deploy and stow. • The EIU has a inhibit relay that controls the power

supply signal from the EEC to the DCV. • The flight crew have control of reverse thrust power

selection. Deploy Pulling up on the thrust reverse select lever in the flight deck will send a signal for thrust reverse select to the EEC and EIU. This will also put the main throttles in the reverse thrust quadrant. The EEC will look for the following conditions before thrust reverse will be allowed; • The EEC will check that the aircraft is on the ground by

checking the LGCU signal of the aircraft computers. • The EEC will check that the engine is running by means

of a N2 signal. • The EIU will look for the signal from the throttle control

unit for energising of the inhibit relay. • The SEC signal for opening the shut off valve. The hydraulic isolation valve solenoid and the directional control valve solenoid will both be energised for a deploy condition. This will admit high pressure hydraulic fluid to the stow and deploy sides of the thrust reverse system. The lower locking actuators will unlock and the EEC sees

a signal from the proximity sensor of reverser system unlocked. In the flight deck this unlocked condition is identified as an amber coloured REV caption on the EPR indicating gauge. When the translating sleeves have moved to 78% of the full deploy position the amber REV indication will change to a green REV indication. When green REV is indicated the full reverse thrust power is available to the flight crew.

Stow To stow the thrust reverse system the flight crew will return the throttles to the idle detent position and select levers to the down position. This will put the throttles back to the forward thrust quadrant. The DCV solenoid will be de-energised as commanded by the EEC via the EIU inhibit relay. This will leave high pressure hydraulic fluid present on the stow side of the reverser system only. As the translating sleeves move from deploy back to stow the flight deck indication will change from green to amber on the REV indication. When the thrust reverser has reached the fully stowed position the amber REV will go and the EPR gauge will return to normal indication. This will indicate that the thrust reverser is fully stowed and locked.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Thrust Reverser Maintenance The thrust reverser system is precisely adjusted to ensure correct alignment and load sharing between the nacelle components and the engine. The thrust reverser actuation system is rigged to synchronise the positions of the left and right translating sleeves and the hydraulic actuators. Improper thrust reverser system rigging can result in a reduction of the service life and/or damage to the actuation system and thrust reverser components.

Maintenance Actions The following will be discussed in the maintenance section of the thrust reverser system;

• Thrust reverser C ducts opening and closing. Opening AMM ref. 78-32-00-010-010. Closing AMM ref. 78-32-00-410-010.

• Thrust reverser system deactivation for maintenance and flight. Deactivation AMM ref. 78-30-00-040-012. Reactivation AMM ref. 78-30-00-440-012.

• Thrust reverser system operation; Manual deploy AMM ref. 78-32-00-860-010. Manual stow AMM ref. 78-32-00-860-011. Hydraulic deploy AMM ref. 78-32-00-860-012. Hydraulic stow AMM ref. 78-32-00-860-013.

• Thrust reverser system order of rigging procedures.

• Thrust reverser C duct rigging. AMM ref. 78-30-00-820-010.

• Translating sleeve and actuators adjust. AMM ref. 78-42-48-400-010.

• Translating sleeve and actuators rigging. AMM ref. 78-32-43-400-010.

• Thrust reverse system synchronisation flex shaft rigging; Removal AMM ref. 78-32-44-000-010 or 78-32-74-000-010. Installation AMM ref. 78-32-44-400-010 or 78-32-74-400-010.

• Lock proximity switch. Adjustment AMM ref. 78-30-00-820-010. Removal AMM ref. 78-31-15-000-010. Installation AMM ref. 78-31-15-400-010.




THRUST REVERSER SYSTEM MAINTENANCE DE

TV25

0263

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Thrust Reverser System Deactivation for Maintenance and for Flight An inoperative thrust reverser may be locked in the forward thrust position for flight. This is as advised by the MMEL requirements for single thrust reverser operation. Thrust Reverser Deactivation for Maintenance Warning Do not cause a blockage of the hydraulic control unit (HCU) return port to deactivate the HCU. If you cause a blockage of the HCU return port the thrust reverser can operate accidentally causing injury or damage. Engine components can stay hot for up to one hour after shut down. Be aware of this when working on the engine immediately after shut down. HCU Deactivation (AMM 78-30-00-040-012) • Carry out the flight deck checks as per aircraft

preparation. • Open the fan cowl doors (71-13-00-010-010). • Position the lock lever on the HCU to the lockout

position and install the deactivation pin. • Ensure that the red pennant is visible to others during

the lockout period. HCU Reactivation (AMM 78-30-00-440-012) • Remove the lockout pin and return the lockout lever to

the usual position. • Close the fan cowl doors (71-13-00-410-010). • Return the aircraft back to the usual condition.

In addition to the procedure for deactivation for maintenance the thrust reverser system can be locked out for flight.

Thrust Reverser Deactivation for Flight In addition to the deactivation procedure for the HCU the translating sleeve can be secured in the stow position by inserting a lockout pin through each translating sleeve and the fixed section of the C duct assembly.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Thrust Reverser C Ducts Maintenance Warning The opening and closing procedure for the thrust reverser C ducts must be adhered to fully. These units can close very quickly and neglect can cause injury to personnel.

Thrust Reverser Opening (AMM ref. 78-32-00-010-010) • Carry out the flight deck checks as per aircraft

preparation.

• Ensure that the area around the engine is clear of obstacles.

• Open the fan cowl doors (71-13-00-010-010).

• Deactivate the HCU (78-30-00-040-012).

• Open the latch access panel and engage the auxiliary latch and take up the tension of the two thrust reverser halves.

• Release the latches in order of; 3, 2, 5, 4, 1.

• Remove the auxiliary latch.

• Attach the hand pump and extend the thrust reverser C ducts to the open position.

• Engage the rear then the front support struts in position and then decay the hydraulic pressure to rest the units on the support struts.

• Disconnect the hydraulic hand pump.

Thrust Reverser Closing (AMM ref. 78-32-00-410-010) • Carry out the flight deck checks as per aircraft

preparation.

• Engage the hand pump and open the thrust reverser C ducts.

• Disengage the support struts and stow them.

• Allow the thrust reverser units to close.

Note: The forward most latch must be in the locked position before closing.

• Engage the auxiliary latch assembly and draw the thrust reverser units together.

• Check front latch has not fouled.

• Disengage the hand pump and engage all latches and lock them in the following sequence; 1, 4, 5, 2, 3.

• Ensure latch unlock indicators are engaged.

• Disconnect auxiliary latch and stow.

• Close the thrust reverser access panel.

• Close the fan cowl doors (71-13-00-410-010).






•

Thrust Reverser System Operation

The thrust reverser can be deployed and stowed manually for maintenance and trouble-shooting operations.

Manual Deploy (AMM ref. 78-32-00-860-010)

The procedure is as follows;

• Open and tag the following circuit breakers for the appropriate engine.

• Open the left and right hand fan cowls.

• Move the thrust reverser hydraulic control unit de-activation lever to the deactivated position and insert lockout pin.

• Disengage the locks on the two locking (lower) actuators and insert pins to ensure locks remain disengaged.

• Position the non return valve in the hydraulic return line to the by pass position.

• Insert 3/8 inch square drive speed brace into external socket and rotate speed brace to deploy the translating sleeve as required.

Note:

Deactivate the system if maintenance is necessary.

Do not exceed maximum indicated torque loading during manual operation.

Deactivate the system if maintenance is necessary. This is done in order to maintain safety during maintenance activities.

Manual Stow (AMM ref. 78-32-00-860-011)

Open and tag the following circuit breakers for the appropriate engine.

• Open the left and right hand fan cowls.

• Move the thrust reverser hydraulic control unit de-activation lever to the deactivated position and insert lockout pin.

• Disengage the locks on the two locking (lower) actuators and insert pins to ensure locks remain disengaged.

• Position the non return valve in the hydraulic return line to the by pass position.

• Insert 3/8 inch square drive speed brace into external socket and rotate speed brace to Stow the translating sleeve as required.

Return the aircraft and engine back to its usual condition.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Thrust Reverser Power Deploy and Stow The thrust reverser system can be operated by using the hydraulic system for control. This is possible with both the engines running and shut down. The advantage of the operation of the thrust reverser system with the engine shut down is to conserve engine life.

Power Deploy • Refer to (AMM ref. 78-32-00-860-012) for this

procedure.

Power Stow • Refer to (AMM ref. 78-32-00-860-013) for this

procedure.

Note: Do not deploy the thrust reverser translating sleeve while the thrust reverser C ducts are open. Damage to the synchronisation cables and the hinged access panels can occur. Be aware of the dangers surrounding the area of the thrust reverser while operating the unit.

If maintenance is to be carried out with the translating sleeves in the deploy position then the thrust reverser system must be deactivated for maintenance. Remove the hinged access doors (HAD) if thrust reverser C ducts are required to be open when the translating sleeves are deployed.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Thrust Reverser System Order of Rigging The thrust reverser system is precisely adjusted to ensure correct alignment and load sharing between the nacelle components and the engine. The thrust reverser actuation system is rigged to synchronise the positions of the left and right translating sleeves and the hydraulic actuators. Improper thrust reverser system rigging can result in a reduction of the service life and/or damage to the actuation system and thrust reverser components.

Order of Rigging The following is a recommended order of rigging of the thrust reverser system components;

• Thrust reverser latches and bumper rigging.

• Thrust reverser translating sleeve and actuators.

• Thrust reverser actuators locks. The LVDTs are self adjusting. After replacement or disturbance of the LVDTs resetting is by cycling the thrust reverser system. There is also a requirement to check system components for satisfactory operation after maintenance has been carried out.

The following table outlines the requirements;

Engine change Latches Bumpers Compression struts

C duct replacement Translating sleeve & actuators Latches Bumpers and compression struts

Translating sleeve replacement

Translating sleeve & actuators Latches

Actuator replacement Translating sleeve & actuators Translating sleeve aft double latches

Flex shafts and tubes replacement

Actuators and flex shafts

Track liner replacement Translating sleeve & actuators CNA replacement Latches

Bumpers



Revision Page 15-41

2

DE

TV25

0263

THRUST REVERSER MAINTENANCE

© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Lock Proximity Switch AMM Ref. 78-30-00-820-010 The lock proximity switch gives indication to the EEC of the thrust reverser lock or unlock condition. In order to maintain the correct function of the switch the distance between the target and proximity sensor must be within the AMM recommendations.

Lock Proximity Switch Check Stow the translating sleeve and then measure the gap between the target and proximity sensor to the AMM recommendations. If the measurements are out of limit then an adustment is necessary.

Rig the Lock Proximity Switch From the measurements taken during the check a spacer is required to adjust the setting to within the AMM recommendations. Disconnect the sensor and target assemblies. Select the correct spacer for the target. Grinding of the selected spacer may be required so as to achieve a greater accuracy for the setting of the target and sensor.




© IAE International Aero Engines AG 2000 IAE V2500 Line and Base Maintenance Thrust Reverser System Thrust Reverser HarnessThe thrust reverser harness connections are shown below.




SECTION 16

TROUBLESHOOTING

© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting IAE V2500 Troubleshooting Introduction In order to locate the source of an engine problem both quickly and efficiently, it is essential that the aircraft maintenance engineer is aware of the fundamental approach to troubleshooting required on the Airbus A319/320/321. Having acquired the knowledge of various engine systems functionality and operation during this course, we are now in a position to take the course the natural step forward and discuss the all-important methodology of isolating and identifying the source of a problem. An important tool available to the engineer is the A319/A320/A321 ‘Computer Assisted Aircraft Trouble Shooting’ (CAATS) CD-ROM. This valuable aid provides the user with an enormous amount of detail and information. This manual is revised and issued every three months. Upon receipt and installation of the up-dated version the previous version is automatically overwritten and the previous disc is now no longer valid or useable. The CAATS CD-ROM is password protected for each airline, as it is tailored specifically for each operator’s requirements. Therefore it is essential that the user always access the procedures for their own particular airline. After inserting the correct password the user is presented with the screen shot below.

The most common and straightforward menu selection is ‘Trouble Shooting Procedures’. If the user were to select ‘Trouble Shooting Manual’, this would require the user to insert a known trouble shooting task reference number in order to progress. Unless a procedural task has been already identified during previous investigation activity it will not be practical to use, as its selection would be dependent knowing which system has the fault. Below is a screen shot showing the opening menu options and for the purpose of these training notes we are using ‘British Midland’ as the ‘Log on Airline’.


© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting


© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting IAE V2500 Troubleshooting Introduction continued: Upon selecting ‘Trouble Shooting Procedures’ the user is then presented with the screen shot (fig. 1) shown below.

From this menu it is possible to enter into the trouble shooting process with information derived from a variety of sources: 1. ECAM WARNINGS: These are the messages that appeared on the Upper ECAM Screen during operation and show the symptom or system, which has been degraded by a fault.

These are generally of a ‘Class 1’ level which would prevent the aircraft from being dispatched unless the problem and source of the message had been rectified. Check Minimum Equipment List (MEL). 2. ECAM STATUS (Inoperative Systems and

Maintenance Status). The presence of an ECAM Status Message ‘STS’ is automatically displayed on the Upper ECAM Screen during Flight phase 1 (Electrical Power ‘on’ before first engine start) and Flight Phase 10 (When the second engine has been shut down after the flight). It is used to highlight a problem or degradation in the built in redundancy facility of the FADEC System. This feature prevents un-wanted distractions of system degradation being shown to the pilot during the flight. A fault of this nature is dispatchable and the fault can be left un-rectified for up to ten days. Check Minimum Equipment List (MEL) the Status Page can then be selected by pressing the STS button on the Systems Page Select Panel. This will then provide information under the ‘Maintenance’ heading regarding the failure, for example ENG 1(2) FADEC or ENG 1 (2) EIU. Fig.1 3. LOCAL WARNINGS (Panel Lights and Standby

Indicators). Lists the entry into the Trouble Shooting procedure. Given indicated engine related faults. (This has limited use). Valves’ and Anti-Ice Valve problems.



STATUS INOP SYSTEM

G+B HYD CAT 3 G RSVR L+R AIL SPLR 1+3+5 L ELEV AP 1+2 ENG 1 REV NORM BREAK NW STEER MAINTENANCE APU AIR COND ENG 1 FADEC

MAX SPD…………………….250/.85 APPR PROC DUAL HYD LO PR -IF BLUE OVHT OUT: -BLUE ELEC PUMP….. ON -LG………………………..GRVTY EXTN -LDG SPD INCRMT………10 KT SLATS SLOW CAT 1 ONLY CANCELLED CAUTION NAV IR 2 FAULT PSI 35 Class 2 Failure of the

Engine Number One’s FADEC System

TAT - 5°C SAT – 30°C

G.W. 60300KG23H56

STATUS PAGE (LOWER ECAM) Revision 2 Page 16-4

© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting IAE V2500 Troubleshooting Introduction continued: 4. FLAGS and ADVISORIES (On ECAM and EFIS

System Pages) Selecting this provides the user with the screen shown Below, see (fig. 2).

By selecting the appropriate system, the user will be presented a complete listing of Flags and Advisories available, related to problems with that particular system. In this example the Engine System. See (fig 3) below.

(fig. 2) (fig.3)




© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting IAE V2500 Troubleshooting Introduction continued: 5. CREW and MAINTENANCE OBSERVATIONS By selecting this option, the user can relate to what conditions they have seen during the engines operation and link and match the symptoms that to the list provides. Fig 4 below illustrates that if the user types in the main heading for that system, a complete list of all possible observations of faults are produced.

Fig. 5 illustrates the complete listing of, in this example (ATA 73) referenced observations

Fig 5 Fig 4




© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting IAE V2500 Troubleshooting Introduction continued: 6. CFDS FAULT MESSAGE Centralised Fault Display System (CFDS) This menu selection is one of the most common methods of entering into the Trouble Shooting process. By interpreting the information provided on the Post Flight Report (PFR) and completing the necessary data field boxes, the user can quickly locate the appropriate Trouble Shooting task for this particular systems problem. In this example we have a problem with the Number 2 Engine’s Fuel Heat Management System. This message appeared on the upper ECAM as an ‘ECAM WARNING’ this is a Class 1 failure and is not dispatchable. The CFDS Fault Message is the text contained under the heading ‘FAILURE MESSAGES’ on the Post Flight Report. Again, in this example the Failure Message that is linked to the Upper ECAM ‘ENG 2 FUEL HEAT SYS’ is:

‘FUL DIV RET VLV/HC/EEC2’

This text along with the ATA reference number:

‘73-13-42’ and the Source:

‘EIU2FAD’ is copied into the text boxes as shown on (fig. 6) opposite.

Completing the Class of Failure data can make further refinement of identifying the task. In this case we had an upper ECAM warning message, so in this example we can identify it as a ‘Class 1’ fault.

Fig. 6






IAE V2500 Troubleshooting Introduction continued: 7. None In selecting the menu option of ‘None’ the user is presented with the screen shown below (fig. 7) This requires the insertion of known information in order to refine the search. If in the example shown the user simply types the first two ‘ATA’ digits for engine related problems, which are ‘77’ and then selects ‘Enter’. Then the complete list of failures and associated warnings is produced (fig. 8) opposite.

Fig 7 Fig 8

© IAE International Aero Engines AG 2000


V2500 Maintenance Special Troubleshooting

Beginning of PFR recording, first engine start + 3 minutes =18:27 End of PFR Recording. 80 knots + 30 seconds = 21:17

GMT (Greenwich Mean Time) =Time when the cockpit warning was displayed. PH = Flight Phase. ATA = Air Transport Association

Note time of ECAM Warning and CFDS Failure Message is the same. (Although there can be up to two minutes difference)

Source = System detecting the fault

POST FLIGHT REPORT

© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting E V2500 Troubleshooting Centralised Fault Display System (CFDS) The purpose of the CFDS is to give the maintenance engineers a central maintenance aid to intervene at system or sub-system level from a Multipurpose Centralised Display Unit (MCDU) located on the flight deck. The MCDU allows the engineer to;

• Interrogate a variety of systems using Built in Test Equipment (BITES) for maintenance information.

• To initiate system return to service tests. The detection of the failures, processing and formatting of the failure messages to be displayed is carried out in each systems individual systems BITE. There are two MCDU’s and either the Captain or First Officer’s MCDU can be used.



Revision 2

Alpha Keys

Line Select Keys

Numeric Keys

Brightness Adjust

Annunciators

Function and Mode Keys

Line Select Keys

Page 16-14

MULTIPURPOSE CENTRALISED DISPLAY UNIT (MCDU)

© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting Failure Classification and Master Minimum Equipment List (MMEL) The MMEL cannot be used as a Minimum Equipment List (MEL) due to the fact that it is not related to operational requirements, specific operations or airlines particular definitions. The MMEL can be used as a basis for particular operators own MEL. The MEL should be used to establish dispatchability for a particular operation. The MEL does not include those items that are obviously required for aircraft safety, such as wings, engines etc. The MEL does not include those items that do not affect the airworthiness of the aircraft, such as galley equipment, entertainment system etc. Note; All items, which are related to the airworthiness of the aircraft and not included in the list, are automatically required to operational for each flight. MEL Preamble The MEL is intended to permit operation with inoperative items of equipment for a period of time, until repairs can be accomplished at the earliest opportunity. In order to maintain acceptable levels of safety and reliability the MEL establishes limitations on the duration of and conditions for operation with inoperative equipment. When an item of equipment is discovered to be inoperative, it is reported by making an entry into the Aircraft Maintenance Record/Logbook as prescribed by the Federal Aviation Regulations (FAR).

The item is then either repaired or may be deferred as per the MEL or other approved means acceptable to the Administrator prior to further operation. MEL conditions and limitations do not relieve the operator from determining that the aircraft is in a condition for safe operation with items of equipment inoperative. Federal Aviation Authority (FAA) Repair Intervals All users of an approved MEL, must effect repairs of inoperative systems or components, deferred in accordance with the MEL, at or prior to the repair times established by the following letter designators:

• Category A: To be repaired within the time interval specified in the remarks column of the operator’s approved MEL.

• Category B: To be repaired within three (3) consecutive calendar days (72 hours), excluding the day the malfunction occurred.

• Category C: To be repaired within ten (10) consecutive calendar days), excluding the day the malfunction occurred.

• Category D: To be repaired within one hundred and twenty (120) days), excluding the day the malfunction occurred.




Printer MCDU’s

LOCATION OF MCDU’S AND PRINTER

© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting Failure Classifications There are three (3) levels of ‘Failure Classifications’ and these are signified by the method of notification of their existence to the Flight Crew or to the Maintenance Engineer during ground operation and testing. Class 1 Failures are indicated, by means of the upper ECAM display or local warnings. Procedures to be followed by the operator to help to ameliorate the problem may also be displayed. Class 2 The operator is informed of a Class 2 failure on the ECAM STATUS page, which only shows the system, affected by the Class 2 failure. A white ‘STS’ symbol appears on the upper ECAM. Class 3 The operator is not informed of Class 3 failures. Class 3 failures are only accessible through the Centralised Fault Display System (CFDS) via the MCDU in ‘menu mode’



CURRENTGROUND FAULTS

FADEC Self TestReverser TestIgnitor TestStart Valve TestP2 T2 Heater Test

500 HOURSClass 3 Faults

UNLIMITEDDESPACTH

DATETIME

ATA CHAPTERCELL NUMBER (1 - 60)

Trouble Shooting Datafrom Stored Faults

Faults storedduring the last leg

Faults storedduring previous 63 legs

FAULT ACRONYM

CELL NUMBERFLIGHT PHASE

FLIGHT LEG

ENGINE PARAMETERSWHEN FAULT RECORDED

'FLIGHT' OPERATION IS DEFINED ASENGINE AT IDLE (PLUS 3 MINUTES)

LAST LEGREPORT

(Screen 7)

PREVIOUS LEGREPORT

GROUNDDATA

(Cells 46 - 60)

FLIGHTDATA

(Cells 1- 45)(Screen 9 &10)

TROUBLESHOOTING(Screen 8)

GROUNDSCANNING

SYSTEM TEST SCHEDULEDMAINT REPORT

CLASS 3(Cells 61-69)

FADEC XXMAIN MENU(Screen 6)

CFDS (Screen 2)

FADEC XX (Screen 6)

ENG (Screen

SYSTEM REPORT TEST

MCDU Main Menu (Screen 1)

Clear Language Message


MCDU SCREEN ROUTEMAP





FAULT CLASS FLIGHT CREW ALERT DISPATCH CONDITIONS

ACTION REQUIRED

Class 1

Visual and Audible Warning (Upper ECAM)

NO GO

or GO IF

or GO

Refer to MEL for details

Class 2

Visual Indication (‘STS’ appears on Upper ECAM) Specific Details (Lower ECAM)

GO

Fault must be recorded and

repaired as per MEL

SMR Scheduled Maintenance Report

No Indication

(CFDS must be interrogated

for details)

No Conditions

Repair at next

'A' Check / 500 hours

Class 3

No Indication

(CFDS must be interrogated

for details)

No Conditions

Time Unlimited Fault

(Should be repaired at

earliest convenient opportunity)

FAULT CLASSIFICATION TABLE AND REQUIREMENTS


© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting Operational Test of the FADEC on the Ground Reference Task Number (73-22-00-710-040) Reason for the Job: Use this test to do a check of the FADEC. Note! : Make sure that the power supply to the FADEC has been supplied for a minimum of 30 seconds, whilst still in menu mode before you start the test. Note! : If failures are found during the test, the message ‘SEE GROUND SCANNING MENU’ comes into view on the MCDU. You must then go into the GROUND SCANNING menu of the FADEC and carryout the related trouble shooting. Note! : If the test is to be repeated on the same or alternate channel, you have to go back to the SYSTEM/REPORT TEST menu and wait 30 seconds before you try to carryout the test again. Should you fail to wait the required 30 seconds, upon completion of the test the message ‘NO RUN’ will appear adjacent to the ‘INPUIT/INT. TEST.




< FMGC Before entering CFDS, ensure that the FADEC power supply is ‘on’

< ACARS

< CFDS

< AIDS

TROUBLESHOOTING WITH CFDS


•

•

•

Operational Test of the Thrust Reverse System with the CFDS Ref Task Number (78-31-00-710-041) Reason for the test: Use this test to carryout a check of the Thrust Reverser System operation. WARNING! MAKE SURE THAT:

THE TRAVEL RANGES OF THE THRUST REVERSERS OF ENGINE 1(2) ARE CLEAR OF ALL TOOLS, EQUIPMENT AND PERSONS. THE THRUST REVERSERS ARE CLOSED AND LOCKED THE THROTTLE CONTROL LEVERS OF ENGINE 1 (2) IS IN THE IDLE POSITION (ZERO ON THE SCALE)

Opposite is an extract from the AMM with some relevant important information indicated.


© IAE International Aero Engines AG 2000 Troubleshooting V2500 Maintenance Special


THRUST REVERSER RETURN TO SERVICE TEST

© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting Operational Test of the Ignition System with the CFDS Reference task number (74-00-00-710-041) Reason for the Task: Use this test to carryout an aural check of the Ignitor plug operation. Warning! Make sure there is no air pressure supplied at the Starter Valve Inlet. 1. Select ‘IGNITOR TEST’ option from the System Test Menu

‘The ‘IGNITOR TEST’ Menu comes into view’. Ensure that the ‘ENG/MODE’ Switch is in the ‘NORM’ position.

2. Set ENG/MASTER Control Switch to ‘ON’. ‘SWITCH 1 ENABLED’ comes into view. 3. Select ‘TURN ON IGNITOR’

‘IGNITOR 1 ON’ comes into view Make sure ingitor plug ‘A’ of the engine makes a noise at the same time.

4. Select ‘TURN OFF IGNITOR’ ‘TURN ON IGNITOR ‘ comes into view Check ignition stops! 5. Select ‘SWITCH 1 ENABLED

‘SWITCH 2 ENABLED’ now displayed.

6. Select ‘TURN ON IGNITOR’

‘IGNITOR 2 ON’ now displayed Make sure ingitor plug ‘B’ of the engine makes a noise at the same time.

7. Select ‘TURN OFF IGNITOR’ ‘TURN ON IGNITOR’ displayed Check ignition stops! 8. On the ENG PNL set the ENG/MASTER control switch to

‘OFF’ 9. Push the Left Line Key adjacent to the ‘RESELECT

MASTER LEV OFF’ indication. The ‘SYSTEM TEST’ menu comes into view.

Note! : The ‘RETURN’ indication does not show. To close the test page, use the line key that normally has the return function. This key stays valid. Do the procedure again for channel ‘B’ of the FADEC.




IGNITOR TEST INDICATIONS

© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting Operational Test of the P2/T2 Heater Reference Task Number (73-22-11-710-040)




OPERATIONAL TEST OF THE P2/T2 PROBE HEATER

© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting Operational Test of the Pneumatic Starter Valve with the CFDS Reference Task Number 80-13—51-710-040 Supply the aircraft pneumatic system from a HP ground power or an APU. On the lower ECAM display, make sure that the available air pressure is between 30 psi (2.07 bar) and 40 psi (2.75 bar) Caution: Make sure that the ENG/MASTER 1 (2) Control Switch (On the panel 115vu) is set to off before you start the fuel pumps. Do not run the engine if the fuel inlet pressure is not positive (The fuel pressure is necessary to lubricate the engine fuel pump and the FMU and thus prevent damage).




OPERATION OF THE PNEUMATIC STARTER MOTOR

© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting Trouble Shooting FADEC Faults and Failure types The majority of FADEC failures take the form of the following Acronym ending. These assist in describing the fault. When attached to the end of the abbreviation of the system that is experiencing the problem it is possible to anticipate the troubleshooting process. A complete listing of all FADEC fault acronyms and a description of what they relate to can be found later in this section.

• Track Check Failure (XXXTK).

• Crosscheck Failure (XX.XCF).

• Wrap-Around Failure (XXXWAF).

• Input Latched Failed (XXXL). Track Check Failure (XXXTK) Failure of a system to follow the commands of the EEC within a specified time. The EEC compares the input (positional feedback provided by the LVDT) against commanded position from the EEC. For example;

Clear Language Message (CLM) ATA ACRONYM

VSVA ACT/HC/EEC# 753241 SVATK Stator Vane Actuator ‘Track Check Fault’.

Crosscheck Failure (XX.XCF) A detected difference in the feedback from the sensors of Channel A and Channel B e.g. LVDT’s, thermocouples or micro-switches. The EEC compares the input (positional feedback signal) from Channel A to that of the input from Channel B. This is only carried out on a EEC input circuit. Example:

Clear Language Message (CLM) ATA ACRONYM VSVA ACT/HC/EEC# 753241 SVAXCF

Stator Vane Actuator ‘Crosscheck Failure’.


© IAE International Aero Engines AG 2000 V2500 Maintenance Special Troubleshooting Trouble Shooting FADEC Faults and Failure types Input Latch Failed (XXXL) A detected failure of an input of the system. The EEC checks the input signal from a feedback device for range and rate of change. This test is only carried out on the input signal to the EEC. Clear Language Message (CLM) ATA ACRONYM

VSVA ACT/HC/EEC# 753241 SVAL Stator Vane Actuator ‘Input Latch Failed’. Wraparound Failure (XXXWAF) A detected failure in the circuitry of a system. The EEC checks the system for continuity. This test is only carried out on an EEC output circuit. Note; The devices that are associated with wraparound faults are solenoids, torque motor windings and micro-switches. Clear Language Message (CLM) ATA ACRONYM

VSVA ACT/HC/EEC# 753241 SVAWAF Stator Vane Actuator ‘Wraparound Fault. Note; By definition the failure message will be set in ‘both’ channels i.e. If ‘channel A’ feedback is not equal to ‘channel B’ then by default ‘channel B’ is not equal to ‘channel A’.

Note; For the purpose of identifying the problem, the Channel that is experiencing a fault will have additional fault messages in the respective channel e.g. ‘Latch Input’ or ‘Track Check’ faults. A complete list of fault code acronyms can be found in the CAATS program by selecting ‘Supporting Data’ and after selecting an appropriate aircraft ‘tail’ number for your airline. Typing in the following ATA Reference; 73-00-00-301 into the ‘Type Known Data’ boxes. You will then be able to view a description of over 250 Fault Code Acronyms.




Accessing Fault Code Acronym Descriptions in CAATS

© IAE International Aero Engines AG 2000 Troubleshooting


V2500 Maintenance Special

Example of Fault Code Acronym Descriptions Contained in the CAATS Program

Documents

V2500 lbm issue 01