V2500 lbm issue 01

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 ABBREIVATIONS
ACAC 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

© 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:
----------------------------------------------------------------------------
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IAE International Aero Engines AG (IAE)
On March 11, 1983, five of the worlds leading aerospace
manufacturers signed a 30 year collaboration 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.
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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.
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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 in
Service
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 for
corporate jet A319 application with increased ‘climb’ rate faciltiy.

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.
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Engine
The V2500 is a twin spool, axial flow, and high bypass
ratio turbofan type engine.
The engine incorporates several advanced 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.
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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

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.
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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.
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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.
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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.
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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).
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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.

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.

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.

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.

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.

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.

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.

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.

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
preparation.
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
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.

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.

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).

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.
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Engine Combined Drains System
Purpose
To provide an early indication of a system or component
failure by evidence of a fluid leak.
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.

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.

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 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.

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.

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.

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.

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.

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.

Nacelle Temperature Sensing and Fire Detection Harness
The nacelle temperature sensing and fire detection
harness electrical connections are shown below.

SECTION 3
MECHANICAL ARRANGEMENT
(Chapter 72)

IAE V2500 Line and Base Maintenance Mechanical Arrangement
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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.
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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.
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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.
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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.
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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.
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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.

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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.
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.
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.

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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

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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.

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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

•
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•
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.

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.

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.

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.

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.

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•
•
•
•
•
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.

•
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•
•
•
•
•
•
•
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).

•
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•
•
•
•
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•
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•
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.

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.

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.

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.

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

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 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.
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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).
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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.
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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

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.

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.

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.

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

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.

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.

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)

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.

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.

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.

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.

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.

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.

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.

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)
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SECTION 5
ELECTRONIC ENGINE CONTROL

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.
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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.
.
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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
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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.
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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).
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Electronic Engine Control (EEC) Data Entry Plug
Purpose
The Data Entry Plug (DEP) provides discrete data inputs
to the EEC. Located on to Junction 6 of the EEC. 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)

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.

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

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

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

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 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
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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.

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.

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.

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.

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.

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.

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 V2500 Line and Base Maintenance Fuel System
Page 7-1Revision 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

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;
-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

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

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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

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

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

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)

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.

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.

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.

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.

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.

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

Fuel SystemIAE V2500 Line and Base Maintenance
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.

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
Fuel flow transmitter
Removal 73-31-17-020-051
Fuel metering unit
Removal 73-22-52-020-051
Fuel distribution valve
Removal 73-13-43-020-010
Fuel distribution valve filter
Removal 73-13-43-000-011
Fuel spray nozzle
Removal 73-13-41-000-010

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.

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.

Fuel Flow Transmitter Removal/Installation
Removal Ch 73-31-17-020-051
• Remove electrical connectors and tubes from the
transmitter.
• Remove the bonding lead.
• Remove the fuel flow transmitter.
• 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).

Fuel Metering Unit Removal/Installation
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).
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).

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.
• 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.
Manual.
• Do an idle check (71-00-00710-012) or wet motor leak
check (71-00-00-710-046).
Fuel Distribution Valve Filter Removal/Installation
Removal Ch 73-13-43-000-011
• Remove the fuel inlet line and the filter cover.
• Remove the filter.
• Install filter
• Install the fuel inlet line and filter cover.
Manual.
Do an idle check (71-00-00710-012) or wet motor leak check
(71-00-00-710-046).

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.
Removal Ch 73-13-41-000-010
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
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.

Fuel System Control Harness
The fuel system control harness electrical connections are
shown below.

Fuel System Harness
The fuel system harness electrical connections are shown
below.

SECTION 8
ENGINE OIL SYSTEM
(Chapter 79)

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

Revision 2 Page 8-2

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
Less than 5 quarts flashes green.
Oil pressure
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.
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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.
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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.
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Rear Bearing Compartment (Bearing no.5)
Purpose
Bearings require oil for:
• Lubrication.
• Cooling.
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.
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• Oil jets directing the oil onto the gears.
High speed external gearbox
Purpose
Gears require oil for:
• Lubrication.
• Cooling.
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

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.

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.

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.

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.

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.

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.
.

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.

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 Engine Oil System
se Maintenance
NO.4 BEARING SCAVENGE VALVE
COOLED HPC12 AIR
IN (BUFFER AIR)
LOW FLOW
HIGH POWER
NO.4 BEARING
SCAVENGE VALVE
OUTLET TO DE OILER
HIGH FLOW
LOW POWER
HPC10 SERVO
AIR IN
NO4 BEARING SCAVENGE OIL
AND VENT AIR OUT TO THE N0.4
BEARING SCAVENGE VALVE.
NO.4 BEARING
SCAVENGE VALVE
DETV250373
VALVE POSITIONAL
FEEDBACK TO EIU

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.

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.

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.

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.

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.

Engine Oil System
PRESSURE TRANSMITTER AND LOW PRESSURE SWITCH
IAE V2500 Line and Base Maintenance
DETV252106

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
Scavenge Filter
Removal AMM ref. 79-22-44-000-010
Oil Scavenge Pump
Removal AMM ref. 79-22-43-000-010
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

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.

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.
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.

Pressure Filter Removal/Installation
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.
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:
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.

Pressure Pump/Filter Housing Removal/Installation
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.
• Install new packings.
• Install sleeve and suction strainer.
• Install the oil pump and filter assembly.
• Connect the air cooled oil cooler supply tube.
Maintenance Manual.
• Fill the engine oil system as necessary (12-13-79-610-
011).
motor leak check (71-00-00-710-045).
•

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.
• 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.
Maintenance Manual.
• Install drain plug.
motor leak check (71-00-00-710-045).
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.

Oil Scavenge Pump Removal/Installation
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.
• 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).

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
In service consumption A1
Fleet range
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

The oil system harness electrical connectors are shown
below.

SECTION 9
HEAT MANAGEMENT SYSTEM

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.
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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.
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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).
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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.
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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:
immediately after the engine is shutdown. the engine
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
• 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

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.

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

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.

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 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.

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.

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.

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.

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.

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.

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.

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

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.

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.

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.

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.

Handling Bleed Valves - Location
The diagram below shows the location of the four bleed
valves and the solenoid control valve.

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
certain threshold)
closed
open (if N2C26
below certain
threshold)
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

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.

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.

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.

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).

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 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.

SECONDARY AIR SYSTEMS INTRODUCTION
DETV250280
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

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 11-5Revision 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
DETV250282
ACC ACTUATOR AND MODULATING AIR VALVE
FAN BYPASS
AIR INLET
TORQUE MOTOR
LVDT FEEDBACK
SERVO FUEL SUPPLIESSERVO FUEL SUPPLIES

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
DETV250283
HPT AND LPT DISTRIBUTION MANIFOLDS

Page 11-9Revision 2
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.

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.

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.

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.

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.

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.

Page 11-21Revision 2
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.

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.

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
DETV250376
CENTRE BEARING
COMPARTMENT COOLING
JACKET
AIR COOLED AIR COOLER (ACAC)

Active Clearance Control System Harness
The HPT and LPT ACC system harness electrical
connections are shown below.

Miscellaneous Systems Harness
The miscellaneous systems harness electrical systems are
shown below.

SECTION 12
ENGINE ICE PROTECTION SYSTEM
(Chapter 30)

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.

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).

Engine Ice Protection SystemIAE V2500 Line and Base Maintenance
Revision Page 12-4
ANTI ICE VALVE
FILTER
LOCKOUT PIN
ELECTRICAL
CONNECTOR
VALVE BODY
ENGINE ANTI ICE VALVEENGINE ANTI ICE VALVE
2
DETV250273DETV250273

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.

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.

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

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
DETV250277
FADEC POWER
SELECT SWITCH
ENGINE START
CONTROL PANEL HAND TUR
POINT
NING
HAND TURNING
POINT
ANTI ICE
VALVELOCKOUT PIN DEACTIVATION
HOLE

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
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.

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
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 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.

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.

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.

Page 13-6Revision 2
DETV250377
UPPER ECAM CRT DISPLAY

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.

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.

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

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
DETV250380

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

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

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.

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.

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
preparation as advised by the AMM.
• 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.

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.

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.

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.

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.

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.

P2 and T2 Probe Harness .
The P2/T2 harness electrical connections are shown below

Temperature Measurement Harness
The temperature measurement harness electrical connections
are shown below.

SECTION 14
ENGINE STARTING AND IGNITION
(Chapter 80)

IAE V2500 Line and Base Maintenance Engine Starting and Ignition System
Page 14-1Revision 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.

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.

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.

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.

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.

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.

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.

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.

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 14-19Revision 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.

Safety Zones - Jet Wake Hazard Areas
During 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

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 5.0
E 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 ISA
dependant
Oil temperature -40 deg.c -10 deg.c 50 deg.c 156 deg.c
amber
165 deg.c red
ENGINE GROUND OPERATION

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.

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
• 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.

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
• 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.

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;
-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.
-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 procedure
71-00-00-700-015 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

Starter Air Valve Harness
The starter air valve harness connections are shown
below.

Ignition System Harness
The ignition system harness electrical connections are
shown below.

SECTION 15
THRUST REVERSER SYSTEM
(Chapter 78)

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.

DETV250262
V2500 THRUST REVERSER SYSTEMV2500 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.

Page 15-5Revision 2
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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

THRUST REVERSER SYSTEM MAINTENANCE
DETV250263

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)
preparation.
• 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.
• 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.

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)
preparation.
obstacles.
• 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)
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.

•
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.

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.

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 15-412
DETV250263
THRUST REVERSER MAINTENANCE

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.

Thrust Reverser Harness
The thrust reverser harness connections are shown below.

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’.

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)

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)

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 5Fig 4

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

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

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

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)

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

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’

CURRENT
GROUND
FAULTS
FADEC Self Test
Reverser Test
Ignitor Test
Start Valve Test
P2 T2 Heater Test
500 HOURS
Class 3 Faults
UNLIMITED
DESPACTH
DATE
TIME
ATA CHAPTER
CELL NUMBER (1 - 60)
Trouble Shooting Data
from Stored Faults
Faults stored
during the last leg
Faults stored
during previous
63 legs
FAULT ACRONYM
CELL NUMBER
FLIGHT PHASE
FLIGHT LEG
ENGINE PARAMETERS
WHEN FAULT RECORDED
'FLIGHT' OPERATION IS DEFINED AS
ENGINE AT IDLE (PLUS 3 MINUTES)
LAST LEG
REPORT
(Screen 7)
PREVIOUS LEG
REPORT
GROUND
DATA
(Cells 46 - 60)
FLIGHT
DATA
(Cells 1- 45)
(Screen 9 &10)
TROUBLESHOOTING
(Screen 8)
GROUND
SCANNING
SYSTEM TEST SCHEDULED
MAINT REPORT
CLASS 3
(Cells 61-69)
FADEC XX
MAIN 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

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.

TroubleshootingV2500 Maintenance Special
THRUST REVERSER RETURN TO SERVICE TEST

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

Operational Test of the P2/T2 Heater
Reference Task Number (73-22-11-710-040)

OPERATIONAL TEST OF THE P2/T2 PROBE HEATER

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

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:
VSVA ACT/HC/EEC# 753241 SVAXCF
Stator Vane Actuator ‘Crosscheck Failure’.

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.
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.
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

Troubleshooting
V2500 Maintenance Special
Example of Fault Code Acronym Descriptions
Contained in the CAATS Program

V2500 lbm issue 01

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