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.
• Ensure that the area around the engine is clear of
obstacles.
• Ensure aircraft leading edge slats are retracted.
• Remove the HAP at the top of the translating cowl if
the thrust reverser is in the deploy position.
• Open the fan cowl doors (71-13-00-010-010).
• Deactivate the HCU (78-30-00-040-012).
• Open the latch access panel and engage the auxiliary
latch and take up the tension of the two thrust reverser
halves.
• Release the latches in the following sequence; 3, 2 ,5,
4, 1.
• Dis-engage the auxiliary latch.
• Attach the hand pump and extend the thrust reverser
C-ducts to the open position.
• Engage the rear then the front support struts in position
and then decay the hydraulic pressure to rest the units
on the support struts.
• Disconnect the hydraulic hand pump.
Closing AMM ref. 78-32-00-410-010
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

•
•
•
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.

•
•
•
•
•
•
•
•
•
HP Turbine
Description
The primary parts of the HP turbine rotor and stator
assembly are:
The HP Turbine Rotor Assemblies (Stage 1 and 2)
The HP Turbine Case and Vane Assembly
The HP turbine rotor assemblies are two stages of turbine
hubs with single-crystal, nickel-alloy blades. The two-hub
configuration removes a bolt flange between hubs. This
decreases the weight and enables faster engine assembly.
The blades have airfoils with high strength and resistance
to creep. Satisfactory blade tip clearances are supplied by
Active Clearance Control (ACC) to cool the case with
compressor air.
The primary parts of the stage 1 rotor assembly are:
Stage 1 Turbine Hub
Inner and Outer HPT Air Seals
64 Blades
Rear HPT Air Seal
The primary parts of the stage 2 rotor assembly are:
Stage 2 Turbine Hub
72 Blades
Stage 2 Blade Retaining Plate
The inner and outer HPT air seals are installed on the front
of the stage 1 hub. The stage 1 blades are installed in
slots on the hub. The blades are held on the forward side
by the outer HPT air seal. The stage 2 HPT air seal is
installed on the rear of the stage 1 hub. This air seal holds
the stage 1 blades on the rear side. Stage 1 blades and
Nozzle Guide Vanes are cooled using H.P. Compressor
discharge air.
The stage 2 turbine hub is installed behind the stage 1 hub
and the stage 2 HPT air seal. Stage 2 blades are installed
in slots in the hub. The blades are held on the forward side
by the stage 2 HPT air seal. The blades are held on the
rear side by the stage 2 blade retaining plate.
Stage 2 HPT blade cooling air is a mixture of HPC
discharge air and stage 10 compressor air. This air passes
through holes in the stage 1 HPT (front inner) air seal and
the stage 1 turbine hub into the area between the hubs.
The air then goes into the stage 2 blade root and out the
trailing-edge cooling holes.

•
•
•
•
•
•
•
•
•
LP Turbine
Description
The primary parts of the Low Pressure Turbine (LPT)
module are:
LPT Five Stage Rotor
LPT Five Stage Stator Vanes
Air Seals
LPT Case
Inner and Outer Duct
LPT Shaft
Turbine Exhaust Case (TEC)
The LP turbine has a five stage rotor, which supplies
power to the LP compressor through the LPT shaft. The
LPT rotor is installed in the LPT case where it is in
alignment with the LPT stators. The LPT case is made
from high-heat resistant nickel alloy and is a one part
welded assembly. To identify the LP turbine module, an
identification plate is attached to the LP turbine case at the
136degrees position.
The LPT case has two borescope inspection ports at
125.27 and 237.10 degrees. The ports are used to
internally examine the adjacent engine sections:
Trailing Edge (TE), Stage 2, HPT Blades
Leading Edge (LE), Stage 3, LPT Blades
The five LPT disks are made from high heat resistant
nickel alloy. The LPT blades are also made from nickel
alloy and are attached to the disks by firtree type roots.
The blades are held in axial position on the disk by the
rotating air seals (knife-edge).

•
•
•
•
•
•
•
•
•
•
•
•
Module 60 - External Gearbox
Purpose
The gearbox assembly transmits power from the engine to
provide drives for the accessories mounted on the gearbox
front and rear faces.
During engine starting the gearbox also transmits power
from the pneumatic starter motor to the core engine.
The gearbox also provides a means of hand cranking the
HP rotor for maintenance operations.
Location
The gearbox is mounted by 4 flexible links to the bottom of
the fan case.
Main gearbox 3 links.
Angle gearbox 1 link.
Description
The external gearbox is a cast aluminium housing that has
the following features;
Individually replaceable drive units.
Magnetic chip detectors.
Main gearbox 2 magnetic chip detectors.
Angle gearbox 1 magnetic chip detector.
The following accessory units are located on the external
gearbox;
Front Face Mount Pads
• De-oiler.
Pneumatic starter.
Dedicated generator.
Hydraulic Pump.
Oil Pressure pump and filter.
Rear Face Mount Pads
• Fuel pumps (and fuel metering unit FMU).
Oil scavenge pumps unit.
Integrated drive generator (IDG).
The Oil sealing for the gearbox to accessory drive links is
provided by a combination of carbon and ‘O’-ring type
seals.
The carbon seals can be replaced while the engine is on
wing.

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.
Revision 2 Page 4-1

Revision 2 Page 4-2

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

Revision 2 Page 4-4

Fan Trim Balance Procedure
Reference 71-00-00-860-010 Engine Operation Limits,
Guidelines and Special Procedures.
Vibration limits and fan trim balance vibration
guidelines:
a) Vibration Limits (Steady State)
− N1 (peak): 5.0 Units.
− N2 (peak): 5.0 Units.
1. Engines that have vibrations within the vibration limits
are acceptable.
2. Engines that have N1 peak vibrations that exceeds the
above limit, troubleshoot as per Trouble Shooting
Manual (ref. TSM task 77-30-00-810-826) or (ref. TSM
task 77-30-00-810-827).
3. Engines that have N2 peak vibrations that exceeds the
above limit, troubleshoot as per Trouble Shooting
Manual (ref. TSM task 77-30-00-810-828) or (ref. TSM
task 77-30-00-810-829).
4. A non-revenue ferry flight to a maintenance base is
permissible with N1 or N2 vibration above limits, if no
fault in the respective trouble shooting procedures in
steps 2 and 3 above. This condition is permissible for
only one engine per aircraft.
b) Vibration guidelines
− N1 (peak): 2.0 Units
1. Fan trim balance is recommended any time N1 peak
vibration exceeds this 2.0 Unit guideline. Perceivable
airframe vibrations generally accompany N1 vibration
levels above this guideline value. Waiting until N1 peak
vibration approaches or exceeds the 5.0 unit limit may
require multiple fan trim balances to bring N1 vibration
down to an acceptable value.
Note:
5.0 units (Aircraft ECAM display) = 1.5 inches per second
of displacement due to the imbalance.
2. Aircraft/Flight Crew Operating Manual (FCOM)
correlation.
− As stated in the FCOM, if N2 vibration during
engine start exceeds limit, the start should be
aborted. Subsequent starts may be initiated without
maintenance action for up to three start attempts.
− The above limits and guidelines are stable (steady
state) and as such may not be stable, therefore the
aircraft level is advisory and not a limit. Vibration
above the advisory level may or may not require
maintenance action, as described in the FCOM;
initially depending on icing conditions or other
engine parameter shifts and finally if the advisory
level is confirmed at steady state conditions.
Revision 2 Page 4-5

Revision 2 Page 4-6

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.

V2500 lbm issue 01

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