V2500 bsi issue 01

V2500 ABBREVIATION
ACAC
ACC
ACOC
AIDRS
Alt
APU
AMM
BDC
BMC
BSBV
CFDIU
CFDS
CL
CNA
C R T
DCU
DCV
DEP
DMC
ECAM
E C S
E E C
Air Cooled Air Cooler E G T Exhaust Gas Temperature
Active Clearance Control
Air Cooled Oil Cooler
Air Data Inertial Reference System
Altitude
Auxiliary Power Unit
Aircraft Maintenance Manual
Bottom Dead Centre 7
Bleed Monitoring Computer
Booster Stage Bleed Valve
Centralised Fault Display Interface Unit
Centralised Fault Display System
Climb
Common Nozzle Assembly
Cathode Ray Tube
Directional Control Unit
Directional Control Valve
Data Entry Plug
Display ManagementComputer
Electronic Centralised Aircraft Monitoring
Environmental Control System
Electronic Engine Control
EHSV
EIU
EIS
EVMS
EVMU
EPR
E T O P S
FADEC
FAV
FCOC
F C U
FDRV
FSN
FMGC
FMV
F M U
F O B
FWC
HCU
HIV
HElU
HP
Electro-hydraulic Servo Valve
Engine Interface Unit
Entered Into Service
Engine Vibration Monitoring System
Engine Vibration Monitoring Unit
Engine Pressure Ratio
Extended Twin Engine Operations
Full Authority Digital Electronic Control
Fan Air Valve
Fuel Cooled Oil Cooler
Flight Control Unit
Fuel Diverter and Return to Tank Valve
Fuel Spray Nozzle
Flight Managementand Guidance Computer
Fuel MeteringValve
Fuel Metering Unit
Fuel On Board
Flight Warning Computer
Hydraulic Control Unit
Hydraulic Isolation Valve
High Energy Ignition Unit (igniter box)
High Pressure

HPC
HPT
HPRV
HT
IDG
IAE
IDG
IFSD
IGV
Ibs.
LE
LGClU
LGCU
LH
LP
LPC
LPCBV
LPSOV
LPT
LRU
LT
LVDT
MCD
MCDU
High PressureCompressor
High PressureTurbine
High Pressure RegulatingValve
HighTension (ignition lead)
Integrated Drive Generator
InternationalAero Engines
Integrated Drive Generator
In-flight Shut Down
Inlet Guide Vane
Pounds
Leading Edge
Landing Gear and Interface Unit
Landing Gear Control Unit
Left Hand
Low Pressure
Low PressureCompressor
Low Pressure Compressor Bleed Valve
Low PressureShut off Valve
Low Pressure Turbine
Line Replaceable Unit
Low Tension
Linear Voltage DifferentialTransformer
Magnetic Chip Detector
MultipurposeControl and Display Unit
MCLB
MCT
Mn
MS
NAC
NGV
NRV
NI
N2
OAT
OGV
OP
OPV
os
Pamb
Pb
PRSOV
PRV
PSI
PSlD
PMA
P2
P2.5
P3
Max Climb
Max Continuous
Mach Number
Micro Switch
Nacelle
Nozzle Guide Vane
Non-ReturnValve
Low Pressure system speed
High Pressure system speed
Outside Air Temperature
Outlet Guide Vane
Open
Over PressureValve
Overspeed
PressureAmbient
Burner Pressure
Pressure Regulating Shut Off Valve
Pressure RegulatingValve
Pounds Per Square Inch
Pounds Per Square Inch Differential
Permanent MagnetAlternator
Pressureof the fan inlet
Pressureof the LP compressoroutlet
Pressureof the HP compressor outlet

P4.9
QAD
SAT
SEC
STS
TA1
TAT
TAP
TCT
TDC
TE
TEC
TFU
TRA
TLA
TLT
TM
TO
Pressureof the LP turbine outlet
Quick AttachlDetach
Static Air Temperature
Spoiler Elevator Computer
Status
Thermal Anti Ice
Throttle Angle Transducer
Transient Acoustic Propagation
Temperature ControllingThermostat
Top Dead Centre
Trailing Edge
Turbine Exhaust Case
Transient Fuel Unit
Throttle ResolverAngle
Throttle LeverAngle
Temperature Limiting Thermostat
Torque Motor
Take-off

INTERNATIONALAERO ENGINES AG V2500 BORESCOPE PRACTICESCONTENTS
Section 1
Section 2 Borescope Requirements.
Section 3 Engine Maintenance Practices.
Section 4 Engine Inspection/Check.
Section 5
Section 6 Typical Examples
Introductionand Engine Mechanical Arrangement.
Part ConditionlClassificationof Damage.

INTRODUCTION AND ENGINE MECHANICAL
ARRANGEMENT

0IAE InternationalAero Engines AG 2000
IAEV2500 Borescope Practices
’ IAE V2500 Borescope Practices
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, hence, periodically. We
welcome suggestions for improvement, and although we
hope there are no errors or serious omissions please let
us know if you notice any.
Telephone: outside UK 00 44 1332 244308
Inside UK 01332244308
Your instructor for this course is:
Introductionand MechanicalArrangement
Initial issue Page 1-1

8 IAE InternationalAero Engines AG 2000
IAEV2500 BorescopePractices Introductionand MechanicalArrangement
IAE InternationalAero Engines AG
On March 11,1983, five of the worlds leading aerospace
manufacturers signed a 30 year collaboration agreement
to producean engine for the single isle aircraft market with
the best proven technology that each could provide. The
five are:
0 Rolls Royce plc-United Kingdom.
0 Pratt and Whitney-USA.
0 Japanese Aero Engines Corp-Japan.
0 MTU-Germany.
0 Fiat Avio-Italy.
Note: Fiat Avio have since withdrawn as a partner.
In December 1983 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 with it's headquarters
in East Hartford, Connecticut,USA.
T o find a name for the engine IAE combined the Roman
numeral V representing the original five partners and the
number 2500 as an abbreviation of the initial engines
maximum thrust of 250001bs.
The V2500 high ratio by-pass turbofan engine to power
the Airbus A320 the 120-180 seat aircraft was launched
on January I"1984.
Each of the share holder companies were given the
responsibility for developing and delivering one of the five
engine modules. They are:
0 Rolls Royce plc--high pressure compressor.
Pratt and Whitney--combustor and high pressure
turbine.
0 JAEC--fan and low pressure compressor.
0 MTU--low pressure turbine.
Fiat Avio--external gearbox.
Note: Rolls Royce have introduced the wide chord fan to
the V2500 enginefamily.
The senior partners Rolls Royce and Pratt and Whitney
assemble the engines. IAE is responsible for the co-
ordination of the manufacture and assembly of the
engines, sales, and marketing and in service support of
the V2500.
Initialissue Page 1-2

IAE V2500 BorescopePractices
IAE V2500 EnginelAirframeApplications
The V2500 engine has been designated the V because
International Aero Engines (IAE) were originally a five
nation consortium. The V being the Roman numeral for
five.
The 2500 numbering indicates the first engine type to be
released into production rated at 250001bs of thrust.
For easy identification for the present and future variants
of the V2500 A5/D5, IAE introduced the following engine
designation system.
All engines will retain the V2500 numbering as a
generic name.
The first three characters of the full designation are
V25. T h i s 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 aircraft manufacturer.
The last figure represents the mechanical standard of
the engine.
The designation V2500-D collectively describes all
applications for the Boeing McDonnell Douglas aircraft.
The V2500-A collectively describes all the applications
for the Airbus lndustrie aircraft.
This is irrespective of engine thrust rating.
The number given after the alpha indicates the mechanical
standard of the engine. For example;
The V2500 Alengine is exempt from these idents as it
was certified with only one thrust rating.
V2528-A5.
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.

IAEV2500 BorescopePractices
0 V2533 A5 -EIS 1997
AIRBUS A321
THRUST 330001bs
V2530 AS-EIS 1994
AIRBUS A321
THRUST 300001bs
V2528 DBEIS 1995
BOEING MD90
THRUST 280001bs
V2527 AS-EIS 1993
AIRBUS A320
THRUST 265001bs
0 V2500 A1-EIS1989
AIRBUS A320
THRUST250001bs
V2524 AS-EIS1997
AIRBUS A319
THRUST240001bs
BOEINGMD90
THRUST250001bs
0 V2525 D5-EIS1995
V2522 ABEIS
AIRBUS A319
THRUST220001bs
33
30
28
26.5
25
22
-_......
*0A-f ......* a... I.. ...............
AIRBUS A321
AIRBUS A320
AIRBUS A319
BOEING MD90
t
0
10
8
E1
IAE V2500 ENGINE/AIRFRAMEAPPLICATIONS t;0

Introductionto 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:
0 The intake cowl.
0 The fan cowl doors.
0 Hinged C ducts with integral thrust reverser units.
0 Common nozzle assembly.
Intake Cowl
The intake cowl allows the smooth intake of air to the
engine while providing an aerodynamic exterior to reduce
engine drag.
The intake cowl contains the minimum of accessories. The
two main accessories that are within the intake cowl are:
0 P2m2 probe.
0 Thermal anti icing ducting and manifold.
Fan Cowl Doors
Access to the units mounted on the fan case and external
gearbox can be easily gained by opening the hinged fan
cowling doors.
The fan cowl doors are hinged to the aircraft strut in four
positions and secured in the closed position by four
latches.
The fan cowl doors have two integral support struts that
are secured to the fan case both are required to hold the
fan cowl doors in the open position.
C Duct Thrust Reverser units
The two C ducts are hinged to the aircraft strut in four
positions and secured in the closed position by six latches
located in five positions.
Each of the C ducts is opened by a hydraulic actuator and
is held in the open position by two integral support struts.
Common Nozzle Assembly (CNA)
The CNA exhausts both the fan stream and core engine
gas flow through a common propulsive nozzle.

.. .....--.
Q IAE lnternatlonal Aero Engines AG 2000
IAE V2500 Borescope Practices Introductionand Mechanical Arrangement
IAEV2500 PROPULSION UNIT

Introduction and Mechanical Arrangement
Engine
The V2500 is a twin spool, axial flow, high bypass ratio
turbofan type engine.
The engine incorporates several advanced technology
features that include:
0 Full Authority Digital Electronic Control (FADEC).
0 Wide chord fan blades.
Single crystal HP turbine blades.
'Powdered Metal' HP turbine discs.
0 A two piece, annular combustion system, which utilises
segmental liners.
Engine MechanicalArrangement
The low pressure (LP) system comprises a single stage
fan linked to a multiple stage booster, has:
0 A5/D5 standard four stages.
0 AI standard three stages.
The fan and booster are axial flow type compressors
driven by a five stage LP turbine.
The booster stage has an annular bleed valve that has
been incorporated to improve starting and handling.
Three bearing assemblies support the LP system. They
are:
0 A single ball type bearing (thrust No 1).
0 Two roller type bearings (support, No 2 & 5).
The HP system comprises of a ten stage axial flow
compressor that is driven by a two stage turbine. The HP
compressor has variable inlet guide vanes (VIGV) and
variable stator vanes (VSV).
0 The A5/D5 standard has one stage of VlGV and three
stages of VSVs.
The A I standard has one stage of VlGV and four
stages of VSVs.
The HP system utilises four bleed air valves designed to
bleed air from the compressors to improve both starting
and engine operation and handlingcharacteristics.
Two bearing assemblies support the HP system. They are:
A single ball type bearing (thrust No 3).
A single roller type bearing, (support No 4).
The combustion system is of an annular open chamber
constructed from two sections.
There are twenty fuel spray nozzles supplying fuel to the
combustor metered according to the setting of the thrust
lever or the thrust management computer via the FADEC
system.

0 IAE InternationalAero EnginesAG 2000
IAE V2500 Borescope Practices Introductionand MechanicalArrangement
COMBUSTOR
HP COMPRESSOR I
BOOSTER
LP COMPRESSOR
AIRCRAFT STRUT
A51D5
0
0
In
ft
2W
0
ACCESSORY GEARBOX
PROPULSIONSYSTEM OUTLINE

Q IAE InternationalAero EnginesAG 2000
IAE V2500 Borescope Practices Introduction and Mechanical Arrangement
Engine
Active clearance control (ACC) turbine
Active clearancecontrol (ACC) is used on both the LP and
HP turbine casings.
This system uses cool air taken from the fan duct.
Engine air bleeds
Engine air bleed is utilised for:
0 Aircraft systems.
0 Compressor stability system.
0 HP and LPturbine active clearance control.
0 10th stage 'make up' cooling air (turbine cooling).
0 Air Cooled Air Cooler ('buffer' air).
0 Air Cooled Oil Cooler.
0 Customer Services Bleed.
HP compressor stage 7 and stage 10 bleeds are available
for aircraft services.
Full authority digital electronic control (FADEC)
The heart of the FADEC is the Electronic Engine Control
(EEC).
The E E C receives rotor speed, pressure and temperature
signals from the engine.
The E E C uses these parameters along with aircraft inputs
to command outputs to engine mounted actuators to
provide control of:
0 Engine fuel flow.
0 Automaticengine starting.
0 Compressor airflow control system.
0 Heat Managementsystem.
0 10th stage make up air system.
0 Thrust reverser.
The E E C also provides protectionfor:
0 N1overspeed.
0 N2 overspeed.
Engine surge.

IAE V2500 Borescope Practices
c
(0 IAE lnternatlonalAero EnginesAG 2OOO
1W
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Initial issue
V2500 ENGINE CUT AWAY
Page 1-10

@ IAE InternationalAero Engines AG 2000
0 Station 3 - HP Compressor exit.
0 Station 4 -Combustion section exit.
0 Station 4.5 - HP Turbine exit.
0 Station 4.9 - LP Turbine exit.
Engine stage numbering
The V2500 engine has compressor blade numbering as
follows;
Stage 1
Stage 1.5
Stage 2
Stage 2.3
Stage 2.5
Stages (3-12)
Engine Leading Particulars For The V2500 A5 Engine
Engine stations
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 the V2500 engine.
The following are the measurement stations for the V2500
engine;
0 Station 1- IntakelEngineinlet interface.
0 Station 2 - Fan inlet.
0 Station 2.5 -LPC OGV exit.
Station 12.5 - Fan exit.
- Fan.
- LPC booster
-LPC booster.
- LPC booster (A5 Only).
- LPC booster.
- HPC Stages.
Note;the HPC is a ten stage compressor.
The V2500 engine has turbine blade numbering as
follows;
Stages (1-2) - HP Turbine Stages.
Stages (3-7) - LP Turbine Stages.
PropulsionUnit Data
T/O thrust (SL static)
Flat ratedtemperature
Total airflow
By pass ratio
Overall pressure ratio
Fan diameter
Propulsionunit length
Engine overall length
Propulsionunit weight
Bare engineweight
250001bs (1I1205kn)
ISA +I5 deg.c
783 Ibs (355kgs)kecond
5.42:l
29.4: 1
63 in (160cm)
198.39 in (503.91cm)
126in(320cm)
73001bs (3311kgs)
49421bs (2242kgs)
Page 1-11Initial issue

0IAE InternationalAero EnginesAG 2000
A I
A51D5
HP COMPRESSOR STAGES
I l l I1.5 2 2.3 2.5
LPC BOOSTERA5/D5
U, /8 LPCSTAGEI
5
tP
ENGINE STATION AND STAGE NUMBERING

0IAE internationalAero EnginesAG 2000
Engine Signals
The following pressure, temperature and rpm signals are
sensed (or derived) by the Electronic Engine Control
(EEC) for power setting systems scheduling and trend
monitoring.
P2 (Fan inlet pressure).
T2 (Fan inlet temperature).
0 P2.5 (LP Compressor Delivery Pressure).
e T2.5 (LP Compressor DeliveryTemperature).
P3 (or Pb) Pressureat the Burner.
e T3 (HP Compressor delivery temperature).
0 P4.9 (or P5) LP Turbine Outlet Pressure.
0 T4.9 (LP Turbine Outlet Temperature).
0 P12.5 (Fan Exit Pressure).
0 N I (Measured) N2 (Derived).
Power
Engine power above idle is controlled and set to an
Engine Pressure Ratio (EPR), which is a ratio of P4.9/P2.
Ternperature
Engine Gas Temperature (EGT) is T4.9.
Trend Monitoring
Trend Monitoring uses signals of P12.5, T2 and T3.
Note: Stations 4 814.5 are not sensed.
Introductionand Mechanical Arrangement

LPTPlSA
SLTO PRESSURE-TEMPERATURE MAP
(A1 engine flgs for 25,0001bsthrust)
126.0 in-
DIMENSIONS
ENGINE DIMENSIONSAND PRESSURE
TEMPERATURE MAP

Module no.
31
Module
Fan
Introduction-ModuleBreakdownof the Engine
The engine is of a modular construction and consists of Modular construction gives the following benefits:
40
50
HP System
LPTurbine
I 32 I IntermediateModule
r 60 I External Gearbox
0 Lower overall maintenance costs.
0 Maximum life achievedfrom each module.
0 Reduced turnround time for engine repair.
0 Reduced spare engine holdings.
0 Easier transportation and storage.
0 Rapid module change with minimum ground running.
0 Easy hot section inspection.
VerticallHorizontal build and strip.
0 Split engine transportation.
Compressor/Turbine independentlybalanced.
Note: Module numbers refer to the ATA section number of
the Chapter/Section/Subject reference in the manuals.
The HPsystem can be further split into mini modules:
0 41 - HP Compressor.
0 42 - Diffuser Case and Outer combustion liner.
0 43 - No 4 Bearing.
0 44 - Stage1 Turbine Nozzle Assembly.
0 45 - HPTurbine

0 IAE International Aero Engines AG 2000
r
31-FAN 32-INTERMEDIATE
60-EXTERNAL
GEARBOX
ENGINE MODULES
initial issue Page 1-16

Module 31-Low Pressure Compressor(LPC)
Purpose
The low pressure compressor is designed to move a large
volume of air rearwards, which constitutes to the majority
of the engines thrust.
Location
The low pressure compressor is attached to the L P C
stubshaft via a curvic coupling. The LPC stubshaft is
attachedto the LPT shaft.
Description
The low pressure compressor is the complete fan
assembly and comprises:
0 22 Hollow fan blades.
0 22 annulus fillers.
0 The fan disc.
0 The front and rear blade retaining rings.
The blades are retained in the disc radially by the dovetail
root with the front and rear blade retaining rings providing
axial retention. Blade removal/replacementis achieved by
removing the front blade retaining ring and the annulus
fillers either side of the blade being removed then sliding
the blade along the dovetail slot in the disc.
In Between the fan blades are annulus fillers forming the
fan inner annulus seal for smooth gas path entry to the
core engine.
The nose cone and fairing smooth the airflow into the fan
but do not form part of the module 31 they are non-
modular parts.

<
0IAE Internatlonal Aero EnginesAG 2000
REAR BLADE
FRONT BLADE
NOSE CONE
LP COMPRESSOR (FAN)L

Module 32 Intermediate Module
Purpose
The intermediate module provides an airflow link between
the L P C booster and HP system. The intermediate module
houses the front bearing compartment and the drive
between gearbox and the core engine.
The fan casing provides a path for the fan to operate in
and also provides protection against a singular fan blade
failure.
Location
The intermediate module is the main mounting location for
the other engine modules.
Description
The Intermediate Module comprises of:
0 Fancase.
0 Fanduct.
0 Fan outlet guide vanes.
0 LPcompressor booster (A5 - 4 stage)(Al - 3 stage).
0 LPcompressor booster bleed valve (LPCBV).
0 Front engine mount structure.
0 Front bearing compartment that houses Nos. 1,2 and
3 bearings.
0 Drive gear for the power off take shaft (gearbox drive).
0 LPstub shaft.
0 Inner support struts.
0 Outer support struts.
0 Vee groove locations for the inner and outer barrels of
the 'C' ducts.

IAEV2500 BorescopePractices Introductionand MechanicalArrangement
Initialissue' '
INTERMEDIATE MODULE FORWARDVIEW
Page 1-20

Module 32 Intermediate Module
Instrumentation
The following pressures and temperatures are sensed and
transmitted to the E.E.C:
P12.5.
0 P2.5.
0 T2.5.
The rear view of the intermediate case i s shown below.
Borescope access to inspect the booster is possible. The
inspection of;
0 Stage 1.5 requires a special tool to guide the flexible
fibrescope through the inlet of the core engine.
0 Stage 2.0 and 2.3 requires the removal of 22 fan
blades, two fan OGV sets and then the removal of a
blank.(A5/D5 engines only)
0 Stage 2.5 access is through the 2.5 booster stage
bleed valve outlet.

INTERMEDIATE MODULE REAR VIEW
initial issue
hl
0
v)
8
E0
Page 1-22

Module 40 HP Compressor
Purpose
The HPC is designed to provide the combustor with high
pressure compressed air.
Location
The HPC is located between the L P C booster and the
combustor.
Description
The HP compressor assembly is a 10 stage axial flow
compressor.
It has a rotor assembly and stator case. The stator case
gives support for the rotor assembly.
The compressor stages are numbered from the front
starting at the L P C booster inlet. This means that the HPC
numbering starts at 3 through to 12.
Airflow through the compressor is controlled by variable
inlet guide vanes (VIGV), variable stator vanes (VSV) and
bleed valves.
A1 standard has:
I stageVIGV.
4 stages VSV.
1stageVIGV.
3 stages VSV.
A5 standard has:
The rotor assembly has five sub-assemblies:
0 Stages 3 to 8 HPcompressor disks.
0 A vortex reducer ring.
0 Stages 9 to 12 HPcompressor disks.
0 The HP compressor shaft.
0 The HPcompressor 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 compressor blades in stages 6 to 12 are installed in
slots around the circumference of the disks. This is
achieved through an axial loading slot, lock blades, lock
nuts and lock screws that hold the blades in position.
The HP compressor stator case has two primary sub-
assemblies:
The HPcompressor front case.
The HP compressor rear case.
All stages of the HPC can be accessed by borescope
inspection. The majority of the blanks, ports and cover
plates are located on the right hand side of the engine.

Introduction and MechanicalArrangement
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8
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Initial issue
'ER
CASE
HP COMPRESSOR ASSEMBLY
Page 1-24

Module40 HP Compressor
Compressor Drums (Rotor)
The rotor assembly is constructed in two parts:
0 The stages 3 to 8 drum.
0 The stages 9 to 12drum.
The two rotor drums are bolted together with a vortex
reducer installed between stages 8 and 9.
The vortex reducer straightens the stage 8 airflow, which
passes to the centre of the engine for internal cooling and
sealing.

introduction and MechanicalArrangement
initial issue
HP COMPRESSOR DRUMS Page 1-26

Module 40 HP Compressor
Compressor Blades
The high pressure compressor is made up of ten stages of
compressor type blades.
0 The first stage is identified as stage 3.
0 The last stage is identified as stage 12.
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 circumferential slot has
two axial loading slots to enable the blades to be installed
into the disc.
Four lock blades are installed on each circumferential slot,
two on each side of the loading slot, which are locked by
lock nuts and lock screws.

STAGES 6 TO 12
BLADES
I dLo
STAGES4 TO
5 BLADES
I
LOCK SCREW
HPC STAGE 3 ’
& LOCK BLADES
AXIAL DOVETAIL
SLOT
CIRCUMFERENTIAL
DOVETAILSLOTS
HP COMPRESSOR BLADES

Q IAE InternationalAero Engines AG 2000
Introductionand MechanicalArrangementIAE V2500 Borescope Practices
Module 40 HP System
Combustor
Purpose
The combustor is designed to mix the compressed air with
fuel and ignitethe mix. The hot gasses are then expanded
to drive the turbines.
Location
The combustor is located between the HPC and HPT.
Description
The combustionsystem is made up of the following:
0 The diffuser section.
The combustion inner and outer liners.
0 The No 4 bearing assembly.
Diffuser Casing
The diffuser section is the primary structural pa
combustion section.
The diffuser section has 20 mounting pads
installation of the fuel spray nozzles. It also
mounting pads for the two ignitor plugs.
Combustion Liner
t of the
for the
)as two
The inner and outer liners form the combustion liner.
Five locating pins that pass through the diffuser casing
locate the outer liner.
The inner combustion liner is attached to the turbine
Initial issue
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.
Page 1-29

/FUEL SPRAY NOZZLE
LOCATING PADS
0
0
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F4
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Initial issue
OUTER COMBUSTION
i LINER
LINER SEGMENTS
COMBUSTOR SYSTEM
JSTION
Page 1-30

Module 40 HP System
Cornbustor
The drawing below shows the arrangement of:
0 The diffuser casing.
The inner and outer combustion liners.
The HPT stage 1NGV’s.
0 The TOBI (Tangential out Board Injector).
Also shown are the No 4 bearing support assembly and
the primary parts of the stage 1turbine nozzle assembly
(HP NGV)
The following components make up the inner liner
assembly of the combustion chamber:
0 The Stage IHPTVane Cluster Assemblies.
0 The Stage 1HPT Cooling DuctAssembly.
0 The Combustion Chamber Inner Liner.
0 The stage 1turbine nozzle assembly that has 40 air
cooled vanes (NGV’s), made of cobalt alloy.
The 40 heat-resistant coated vanes are attached to the
stage 1HPT cooling duct assembly with bolts. The hollow
vane airfoils have internal baffles and cooling holes.

0IAE International Aero Engines AG 2000
HPC OUT1
VANES
COMBUSTIONCHAMBER
OUTER LINER
COMBUSTION
I CHAMBER INNER LINER
FUEL SPRAY
NOZZLE
DIFFUSER CASE //
ASSEMBLY
No4 BEARING /
SUPPORT ASEMBLY
COMBUSTOR CROSS SECTION

Module 40 HP System
HP Turbine (HPT)
Purpose
The HPT is subjected to the passage of the hot expanding
gasses from the combustor, which pass across the aerofoil
surface. This transmits a rotational input to the HPT shaft
and turns the HPC as a result.
Location
The HPT is located between the combustor unit and the
LPT stage 3.
Description
The primary parts of the HP turbine rotor and stator
assembly are:
0 The HP Turbine Rotor Assemblies (Stage 1and Stage
0 The HPTurbine Case and Vane Assembly.
The HP turbine rotor assemblies are two stages of turbine
hubs with single-crystal, nickel-alloy blades with high
strength and resistance to creep. The two-hub
configuration removes a bolt flange between hubs, which
decreases the weight and enables faster engine
assembly.
Satisfactory blade tip clearances are supplied by active
clearance control (ACC) to cool the case with by-pass
airflow.
2).
The primary parts of the stage 1rotor assembly are:
Stage 1turbine hub.
0 Inner and outer HPT air seals.
0 64 Blades.
0 Rear HPT air seal.
The primary parts of the stage 2 rotor assembly are:
0 Stage 2 turbine hub.
0 72 Blades.
0 Stage 2 Blade retaining plate.
Borescope access for the HPT blades is possible for
leading and trailing edges of the stages 1and 2.

IAE Introduction and Mechanical ArrangementV2500 Borescope Practices
HPC loTHSTAGE AIR FOR
HPT STAGE 2
HPT STAGE 1
HP NGV 2
HPT COOLING AIR
FEED TO STAGE 1
HPC 1OTH STAGE AND STAGE 2 HPT
DISCHARGEAIR TO HPT AIRSEAL
Initial issue
HP TURBINE ASSEMBLY Page 1-34

Module 50 LP Turbine
Purpose
The low pressure turbine (LPT)i s designed to maintain the
rotational momentum of the L P C system.
Location
The LP system module is attached to the HP system
module and is linked to the L P C by a singular shaft.
Description
The primary parts of the low pressure turbine (LPT)
module are:
0 LPT Five Stage Rotor.
0 LPT Five Stage Stator Vanes.
0 Air Seals.
0 LPTCase.
0 Inner and Outer Duct.
0 LPTShaft.
0 Turbine Exhaust Case (TEC).
The LP turbine has a five stage rotor that supplies power
to the LPcompressor 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.
The LPT case has two borescope inspection ports found
on the left and right hand sides. The ports are used to
internally examine the adjacentengine sections:
Trailing Edge (TE),Stage 2, HPT Blades.
0 Leading Edge (LE),Stage 3, LPT Blades.
The remaining stages of the LP system do not have
borescope access ports.
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 roots. The blades are held
in axial position on the disk by the rotating air seals (knife
edge).
Stage 3 97 blades.
Stage 4 99 blades.
Stage 5 87 blades.
Stage 6 85 blades.
Stage 7 89 blades.

0 IAE lnternalionalAero EnginesAG 2000
N
E0
Initial issue
LOW PRESSURE TURBINE MODULE Page 1-36

Module 60 External Gearbox
Purpose
The gearbox assembly transmits power from the engine to
provide drives for the accessory units mounted on the
gearbox front and rear faces.
During engine starting the gearbox also transmits power
from the pneumatic starter motor to the 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.
0 Main gearbox 3 links.
0 Angle gearbox 1 link.
Description
The high speed gearbox (HSGB) is a cast aluminium
housingof which the accessory units are mounted onto.
The accessory units receive drive from the gears within
the gearbox by feeder or quill shafts.
The following are the features of the HSGB:
0 Individuallyreplaceable drive units.
0 Magnetic chip detectors.
0 Main gearbox 2 magnetic chip detectors.
0 Angle gearbox 1 magnetic chip detector.
The HSGB has both front and rear faces occupied by
accessory units that serve the function of both the engine
and aircraft. These units are:
Front Face Mount Pads
0 De-oiler.
0 Pneumatic starter.
0 Dedicated generator.
0 Hydraulic Pump.
0 Oil Pressure pump.
Rear Face Mount Pads
0 Fuel pumps (and Fuel metering Unit FMU).
0 Oil scavenge pumps unit.
0 IntegratedDrive Generator (IDG).

OIL FILTER
Tm
-STARTER LOCATION
/HYDRAULICPUMP
LOCATION
OIL TANK
FUEL PUMP DRIVEPAD
OIL SCAVENGEPUMP
PNUEMATIC
OIL PRESSUREPUMP
INTEGRATEDDRIVE
GENERATOR
LOCATION
HIGH SPEED GEARBOX AND ANGLE BOX
TlNG

V2500 BorescopePractices BorescopeEquipment
Borescope Requirements
Borescope equipment permits the inspection of gas
turbine engine parts that would otherwise be inaccessible
with the engine installed and in service.
Engine removal, either due to suspected internal damage
or because of maintenance schedules based on hard time
life philosophy involves high costs to operators. It is an
obvious advantage to allow an engine to remain in service
until one or more of the following reveal defects:
0 Performance analysis.
0 Oil analysis.
0 Borescope inspection.
0 Repetitive monitoring of allowabledamage.
Borescope inspection requirements basically fall into 3
categories:
0 Scheduled inspection.
0 Special inspection.
0 Non scheduled inspection.
Scheduled Inspections
These are regular inspections carried out as part of an
approved maintenance schedule, the frequency of which
is dependantupon either engine cycles or flight times.
'
The combustion and turbine sections are of primary
concern due to the high stresses and temperatures in
these areas. All defects should be recorded, ideally on a
specific chart, to record any deterioration and
assessments can then be made to establish whether the
engine;
0 Continues in service to the next scheduled inspection.
0 Continues in service with reduced periodicitychecks.
0 Is removed either immediately or within a specified
time.
Special inspection
Defects may be highlighted by either service experience or
shop inspection and by the introduction of special
inspections these particular defects can be monitored
whilst the engine remains in service.
Non scheduled inspections
Borescope inspection can be used to great effect to
assess the serviceability of an engine after such incidents
as:
Ingestion of foreign objects.
0 Engine surge.
0 T G T or RPM exceedances.

THIS PAGE IS LEFT INTENTIONALLYBLANK
Initial issue
V2500 Borescope Practices Borescope Equipment
Page2-2

V2500 Borescope Practices
Borescope Equipment
V2500 Borescope Equipment
To effectively carry out a borescope inspection of an
engine minimum equipment requirement is quoted in the
AMM. However there are systems available that are far
more advanced than the minimum equipment levels.
This equipment is complimentary to the engine inspection
and can only improve the inspection above that of the
standard equipment.
IAE Have produced a specification (Part no. IAE6F10408)
which covers the total requirement for performing
inspection on the V2500.
This includes items which can be considered as additional
to the minimum necessary list, such as photographic and
CCTV options.
V2500 Basic Kit
The basic equipment for the successful borescope of the
engine is as follows:
0 High power light source.
0 Light guide cable.
0 Rigid Borescope (5.5 & 8mm diameter).
0 Flexiblefibrescope (6 & 8mm diameter).
0 90"Viewing Adapter.
0 Guide tube - (IAE 2J12030) Stage 1.5 Blade
inspection.
0 Guide tube - (IAE 2P16204) Stage 1 HP Turbine Blade
inspection.
Initial issue
Light Source
The light source concentrates maximum light into the
guide cable.
Standard light source uses a tungsten halogen projector
lamp up to 150watt output.
Due to the large areas of the combustion system, or if
photography/CCTV will be used, a light source using
either metal halide or xenon arc lamp will be necessary,
supplied by 110-240volt AC 50-400Hz power.
For portability some units operate from 12-15 volts DC
sUPPlY
Light Guide Cable
The guide cable connects a light source to either rigid or
flexible borescopes.
In most flexible fibrescopes the guide cable is an integral
part of the unit to reduce light losses.
Page 2-3

Q IAE International Aero Engines AG 2000
Borescope Equipment
STEADY HANDLE

STORAGECASE
io
sa
LIGHTGUIDE CABLE
LIGHTSO
BOROSCOPES
STORAG-ECASE
BORESCOPE INSPECTION EQUIPMENT

V2500 Borescope Equipment
Rigid Borescopes
0IAE internationalAero EnginesAG 2000
BorescopeEquipment
Fibrescopes
For the inspection of the V2500 engine rigid borescopes Fibrescopes are necessary if a full AMM inspection of the
are the best recommendedfor use. engine is to be carried out. Fibrescopes have the
advantage of the reaching areas of the engine that rigids
The following table shows the recommended scopes:
cannot.
The 6mm and 8mm fibrescopes are recommended forDiameter Working length Direction of Field of view
degs use. This recommendation is in conjunction with rigidmm cm view degs
5.5 33 50 35
5.5 33 90 35
5.5 33 115 35
8.0 29 50 50
8.0 28 90 50
8.0 28 115 50
Always inspect the borescope equipment for satisfactory
service prior to using.
Rigid scope identifier alphalnumeric;
F10002400055
F no orbital scan.
G orbital scan.
100 insertion tube diameter.
024 insertion tube working length.
000 direction of view.
55 field of view.
Initial issue
scopes.
Fibrescope identifier alphalnumeric;
IF 804 15
IF industrial fibrescope.
8 nominal diameter.
D direct view.
4 mk4 O E S range.
15 working length.
The interchangeable tip has the same idents
except the working length is replacedwith:
AIOS
A10 field of view.
S direction of view.
s the cope
Page2-5

EYE PIECE
PRO
FOCUS
EYE
000"
045"
090"
110"
ritl;t
ARTICULATION
CONTROL
U
h
Borescope Equipment
DIRECT (BLUE)
FORE-OBLIQUE (GREEN)
LATERAL(RED)
RETRO (YELLOW)
INTERCHANGEABLE
PROBETIPS
FLEXIBLE INSERTION
TUBE
LIGHTGUIDE
RIGID BORESCOPES AND FIBRESCOPES

V2500 BorescopePractices
Care, Use and Storage
Rigid Borescopes
Before use remove protective caps and inspect the
objective and eyepiece windows. If required, clean using
manual procedure:
Lens tissue or cotton applicators.
Lens cleaner alcohol and methylated spirit for any oil
dust or dirt.
0 Check focus control.
Check end faces of light guide cable for cleanness and
damaged areas.
0 Check borescope for damage.
0peration
Insert borescope with care and look in the eyepiece to
check direction of view.
Adjust focus and illumination to give a clear image.
Remove carefully from engine.
Clean if necessary fit protective caps and return to case
for storage.
Initial issue Page2-7

Fibrescopes
Certain V2500 inspections require the use of a fibrescope
of 6mm or 8mm diameter.
Tip movement is limited to two way articulation in the 6mm
unit and four way with the 8mm size.
Before use remove protective caps and inspect the
eyepiece window. If required, clean using manual
procedure:
0 Clean the lens with tissue or cotton applicators.
0 Lens cleaner alcohol and methylated spirit for any oil
dust or dirt.
0 Adjust the diopter ring for a sharp image.
0 Check focus control.
0 Check bending section has full and correct movement.
0 Check the angle free control locks and releases
correctly.
Check borescopeand insertion tube for damage.
Operation.
0 Set angle free control to free position.
Adjust the brightness control on the light source to
obtain optimum illumination.
Confirm direction of view by looking in the eyepiece
and advance the insertion tube slowly without force.
Adjust focus to give a clear image.
After inspection, set the bending section straight, the
angle free control to the free position and remove
carefully from the engine.
Note:
Do not use excessive force to pull the fibrescope from the
engine and support the insertion tube and tip during
removal.

Care of Use of Borescope Equipment
Rigid Scopes
Do
Follow the manufacturers instructions and
recommendations.
Handle with care at all times. Shocks or bends can
damage the optics.
Inspectthe equipment prior to use.
Check the end faces of the light source cable for
clearness and clean as necessary.
Clean the equipment before use and also prior to
stowage before completion.
Always fit protective caps when the scope is not in use.
Do Not
0 Subjectthe instrumentto any unnecessary force.
0 Use in a flammable atmosphere.
0 Bend or hit the shaft.
0 Forcethe focusing barrel against the stops.
0 Forcethe rotational control against the stops.
0 Introducethe scope into live electrical equipment.
0 Totally immerse the scope into liquids (main body).
0 Lay the scope on hard surfaces where it could be
exposed to pressure or weight that could bend the
shaft.

0IAE International Aero EnginesAG 2000
Care of Use of Borescope Equipment
Fibrescopes
Do
Follow the manufacturer instructions and
recommendations.
Handle with care at all times. Shocks or bends can
damage the optics.
Inspect the equipment prior to use.
Clean the equipment before use and also prior to
stowage before completion.
Always fit protectivecaps when the scope is not in use.
Remove any moisture prior to stowage.
Ensure that the o ring seal on the optical adapter is in
good condition and correctly fitted.
Do not
0 Subject the distal end to shocks and impacts.
0 Sharply bend or strain the light guide cable.
0 Sharply bend or strain the insertion tube.
0 Twist the bending section by hand.
0 Insert into live electrical equipment.
0 Use in corrosive fluids.
0 Leave the finger on angle knobs when removing the
scope.
0 Use hard cloth or brush for cleaning.
0 Apply excessive force to the bending section.
0 Apply excessive force when inserting or removing the
insertion tube.

Q.1
Ans.
Q 2
Ans.
Q.3
Ans.
(2.4
What types of borescope are recommended by the
AMM for the V2500 inspection?
a) Rigid borescope only
b) Combination of rigid and flexible borescopes
c) Flexible borescope only
Why are ’special inspections’ performed?
a) T o follow a required maintenance schedule
b) T o monitor damage and its progression
highlighted by shop inspection or service
experience while engine stays in service
assess for continued serviceability
c) T o inspect the engine, following an incident and
What procedure is recommended before replacing a
borescope in a case?
a) Clean and dry the borescope and fit protective
b) Immerse the borescope in cleaning fluid
c) Clean with a stiff brush and hot water
What is the view of a borescope rigid colour coded
‘RED’
a) 1looretro
b) 4 5 O fore-oblique
c) 90° lateral
covers

MaintenancePractices
The safety of personnel and the aircraft is of paramount
importance. The maintenance practice section looks at the
actions that can be carried out which safely allow a
borescope of the engine to take place.
The section looks at the following engine preparations for
borescope:
Aircraft preparation.
Thrust reverse deactivation.
Fan cowl doors opening and closing.
Thrust reverser C ducts opening and closing.
Rotation of the LPand HPshafts.
Igniter plug removal and installation.
Borescope access plugs removal/installation.
Aircraft Preparation
The preparation of the aircraft requires that a check in the
flight deck be carried out.
The check is to ensure that the FADEC system power
switch is in the off position. The switch is located on the
overhead panel 50VU.
Install a DO N O T O P E R A T E identifier on the center
control pedestal on the ENG panel 115VU.
Ensure that the engine has not been shut down within the
last five minutes.

*I.. --. ..
Maintenance Practices
ENGINE CONTROL
PANEL-Vi15
FLIGHT DECK

Thrust Reverser 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. B e aware ofthis when working on the engine
immediately after shut down.
HCU Deactivation(AMM 78-30-00-040-012)
0 Carry out the flight deck checks a s per aircraft
preparation.
0 Open the fan cowl doors (71-13-00-010-010).
0 Position the lock lever on the HCU to the lockout
position and install the deactivation pin.
0 Ensure that the red pennant is visible to others during
the lockout period.
HCU Reactivation (AMM 78-30-00-440-012)
0 Remove the lockout pin and return the lockout lever to
the usual position.
0 Close the fan cowl doors (71-13-00-410-010).
0 Return the aircraft back to the usual condition.

V2500 Borescope Practices Maintenance Practices
HYDRAUL
- HYDRAULIC CONTROL UNIT (HCU) DEACTIVATION ANDU
REACTIVATION

Fan Cowl Doors Maintenance
Warning
Make sure that the landing gear ground safeties and the
wheel chocks are in position.
B e 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 mechanism is
secured.
Fan Cowl Doors Closing
AMM ref. 71-I3-004 0-010
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 71-0259an additional
feature called the hold open device is fitted. T o 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:
SBN 71-0259introduces 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.

V2500 BorescopePractices Maintenance Practices
F

V2500 BorescopePractices MaintenancePractices
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 C Duct Opening
AMM ref. 78-32-00-010-010 ducts.
Thrust Reverser C Duct Closing
AMM ref. 78-32-00410-010
Carry out the flight deck checks as per aircraft
preparation.
Engage the hand pump and open the thrust reverser C
e
e
e
e
e
e
e
e
e
0
Carry out the flight deck checks a s per aircraft Disengage the support struts and stow them.
preparation.
Ensure that the area around the engine is clear of
obstacles.
Open the fan cowl doors (71-13-00-010-010).
Allow the thrust reverser units to close.
Note:
The forward most latch must be in the locked position
before closing.
Deactivatethe 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 1through to 5.
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.
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, and 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.
Initialissue Page3-7

0 IAE international Aero Engines AG 2000
BIFURCATION
HYDRAULIC HAND
s
8
rIn
tn
THRUST REVERSER C DUCT OPENlNGlCLOSlNG

HP System rotation point access
Removalllnstallationaccess cover
The access cover is located on the front face of the
external gearbox between the air starter and dedicated
alternator (AI/A5).
Air starter and hydraulic pump (D5).
T o remove the blank;
Remove the two nuts and washers (Al/A5), two bolts and
washers (05) that secures the cover plate the gearbox.
Remove the cover plate and discard the rubber packing
(seal ring) from the cover.
Turning of the HP Rotor.
The HP system may be turned by using one of the
methods in the AMM.
Manual
Install a 9/16”A F socket to the starter idler gear with a
torque wrench.
Turn the starter idler gear clockwise to turn the HP
compressor in a clockwise direction, or opposite, as
required. Viewed from the front of the engine.
On completion of inspection, remove socket, lubricate with
engine oil and fit new packingto cover.
Refit cover to gearbox and torque load to the specified
values in the manual.
Automatic
Rotator Kit Part No. IAE 2F10057 or 2F10066.
This unit provides a compressed air powered device to
rotate the HP system through the access point on the
gearbox. The unit has foot pedal control of both direction
of rotation and speed of rotation; it also incorporates a
protractorto give an indication of rotation in degrees.
Alternative ElectronicUnits
Various manufacturers have produced sophisticated
electronic rotation controllers with viewing screens,
incorporating methods to tag and return to damage
locations during the inspections, at this time these are not
included in the manual.

HIGH SPEED GEARBOX HAND TURNING
Initial issue
COVER PLATE'
Page 3-10

0 IAE lnternalionalAero EnginesAG 2000
Igniter Plugs Maintenance
Removalllnstallationprocedure
Warning
Do not touch ignition system components for at least one
minute after the ignition power is switched off. Electrical
discharge of the HE unit is dangerous and can kill.
Follow all safety procedures as identified in the Aircraft
MaintenanceManual
Igniter Plug Removal
AMM ref. 74-21-41-000-010
0 Remove clamps and cooling shrouds.
0 Disconnect lead and discard grommet.
0 Disconnect the clip positions that restrict movement of
the lead.
0 Remove the plug, but not the housing installed in the
case. This is adjusted by shim washers and presets
the immersion depth of the Igniter plug at engine build.
Caution
Do not bend the ignition lead too much when you
disconnect it, the lead can be damaged and cause
electrical circuit defects.
Igniter Plug Installation
AMM ref. 74-21-41-400-010
The procedure for installation is the reverse of removal
except for the following points:
Ensure plug threads are clean and not damaged.
0 Torque tighten all locations to the AMM specified
values.
0 After completion perform an Igniter plug function test.

COOLING SHROUD
@
2 IGNITERLEAD
E0 IGNITER PLUG REMOVAL/INSTALLATlON

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
AI 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. This will give access to the trailing edge of stage
2.0 and the leading edge of stage 2.3.

V2500
2.5
LP COMPRESSOR BOOSTER BORESCOPEACCESS
Page 3-14Initialissue

HPC BorescopeAccess
There are two SB standards of HPC designs for A I
engines where the borescope access features differ. The
following is an explanation of the pre and post modification
standards.
A1 Engines pre SB 72-0033
The following information relates to engines that are prior
to SN V0127.
The inspection requirements as advised by the AMM for
this standard of engine is:
0 Recommended that only stage 3 and stage 12 of the
HPC blades be inspected while the engine is on wing.
0 Recommended not to use access port D as possible
damage can be caused to the inspection equipment.
0 A limited number of engines prior to SN V0063 will
have HPC borescope blanks produced to pre SB 72-
0005.
SB 72-0005 introduces a borescope plug of reduced
weight.
SB 72-0033 introduces scalloped heatshield retainer at the
stage 7 (access port D) borescope position.
SB 72-0027 introduces a weight reduced HP compressor
front case assembly.
A1 Engines pre SB 72-0027
The borescope access ports give access to these stages
of the compressor:
Port A HPC stage 3 rear. It is located 34 deg below the
engine horizontal on the right side.
Port B HPC stage 3 rear and stage 4 front. It is located 34
deg below the engine horizontal on the left and right side.
Port C HPC stage 5 rear and stage 6 front. It is located 67
deg below the engine horizontal on the right side.
Port D HPC stage 7 rear and stage 8 front. It is located 67
Port E HPC stage 8 rear and stage 9 front. It is located 74
Port F HPC stage 9 rear and stage 10front. It is located
61 deg below the engine horizontal on the right side.
Port G HPC stage 11rear and stage 12 front. It is located
55 deg below engine horizontal on the right side.
Note:
The access to the borescope plug C requires the removal
of the control rod from the unison ring at the stator six
VSVs. It is recommended that only the stage 3 and stage
12 HPcompressor blades be examined with the engine
on-wing.

PREMODSB 72-0100
PORTSE, F, G
POST MOD SB 72-0100 PORT D POSTMODSB 72-0100
SB 72-0100
PORT D PRE MOD
PORTC
> &
PORTbA
HPC BORESCOPE ACCESS A I PRE SB 7210027

A I Engines post SB 72-0027
of the compressor:
Port A HPC stage 3 rear. It is located 34 deg belowthe
engine horizontal on the right side.
Port B HPC stage 3 rear and stage 4 front. It is located34
deg belowthe engine horizontal on the left and right side.
Port C HPC stage 5 rear and stage 6 front. It is located
33.5 deg below the engine horizontal on the left side.
Port D HPC stage 7 rear and stage 8 front. It is located67
deg belowthe engine horizontal on the right side.
Port F HPC stage 9 rear and stage 10front. It is located
61deg below the engine horizontal on the right side.
Port G HPC stage 11rear and stage 12front. It is located
Port H HPC stage 3 front (additional access). It is located
41 deg below the L.Hsplit line to give a greater inspection
capability.
Note:
It is recommended that only the stage 3 and stage 12 HP
compressor blades be examined with the engine on-wing.
The access to the borescope plug C requires the removal
of the control rod from the unison ring at the stator six
vsvs.
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 recommendedby the AMM.
Take care not to let excessive jointing compound enter the
borescope access port hence into the engine.

. .. .. ... .
PORT H
PORT E F G
C
> B PORTA
PORT B
PORT C
SB 72-0338
PORT D
PORT C
SB 72-0265
3N
0
In
[3 HPC BORESCOPE ACCESS A I POST SB 72-0027
Page 3-18Initialissue

Q IAE International Aero Engines AG 2000
A5/D5 Engines HP Compressor Borescope Access
of the compressor:
Port A HPC stage 3 front. It is located40 deg below the
engine horizontal on the left side.
Port B HPC stage 3 rear and stage 4 front. It is located 34
Port C HPC stage 5 rear and stage 6 front. It is located 61
deg below the engine horizontal on the left side.
Port D HPC stage 7 rear and stage 8 front. It is located 61
Port F HPC stage 9 rear and stage 10 front. It is located
61 deg below the engine horizontal on the right side.
Port G HPC stage 11rear and stage 12front. It is located
Note:
It is recommendedthat only the stage 3 and stage 12 HP
compressor blades be examinedwith the engine on-wing.
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 compoundenter the
borescope access port hence into the engine.
Initial issue
Page3-19

-
PORT E F G
PORTC
PORT C
SE 72-0317
PORTA
B PORT B
PORTC
HPC BORESCOPE ACCESS AWD5

HP CompressorAccess Port C-A1 Engines
The HP Compressor borescope access port C requires
the removal of the stage 6 VSV control rod. This will allow
access to be gained for borescope inspection of the rear
of rotor 5 and the front of rotor 6.

CON1
CENTRALISINGPAD
/ VSV OPERATING
LEVER
STAGE 5 UNISON
RING
BORESCOPE ACCESS
PORT C
’ HP COMPRESSOR BORESCOPE ACCESS PORT C

Combustor, HP and LP Turbines BorescopeAccess
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
A I 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:
0 B1to 84 for the left hand side of the engine.
0 B5 and the 2 igniter plug ports for the right hand side of
the engine.
A I 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:
0 B1to B5 for the left hand side of the engine.
0 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 1the 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 a s recommended by the AMM.
SB 72-0221 introduces a new diffuser case assembly.

Maintenance PracticesV2500 Borescope Practices
<
I
/
8
TRAILING EDGE HPT 1
LEADING EDGE HPT 2COMBUSTION CHAMBER
COMBUSTION
CHAMBER
TRAILING EDGE HPT 2
LEADING EDGE LPT3
IGNITER PLUG
03
d
N
2' COMBUSTION CHAMBER, HPT AND LPT BORESCOPE
ACCESS PORTS
n

Engine Inspection/Check
Borescope Terms of Identification
During borescope inspection of the engine general terms
which identify a condition of damage is used. For example;
Cracks.
Nicks.
Burns.
Tears.
However the AMM ask that certain damage observed be
given a term of descriptionfrom that of the normal.
These terms are as follows:
Combustion Chamber
Burn hole
A local burn area that has continued through the base
material. The ratio of a burn holes circumferential length to
axial length usually will be lessthan or equal to 2.0.
Edge Burnback
Full thickness liner segment burns that start at an edge,
occur more at the segment trailing edge location but have
also occurred at the leading edge corners. Burnback is
material that is not there that has a ratio of the burn areas
circumferential length to axial length of more than 2.0.
Radial Burnback
Full thickness burning on deflector edge continuing radially
inward towards fuel nozzle.
HPT Is'Stage Vane
Burn Through
A local burn area that has continued through the aerofoil
surface. Cracks that have burned too more than 0.05in
(1,27mm) separation are also burn through.
Trailing Edge Burns
Burns or burn through that start at the aerofoil trailing edge
and continueforward to the aerofoil leading edge.
Lift Up
Lift up is where the surface on one side of the crack is
higher than the surface on the other side of the crack.
Coating Damage
Chips, spalling, flakes, blisters, peeling or oxidation that is
confined to the coating.
Connected Loop Crack
A crack or a group of cracks that show an isolated area of
vane metal.

BorescopeTerms of Identification
HPT stages 1 and 2
Erosion
A local area where material has been removed by causes
other than heat stress.
Leading Edge Burn Through
A hole in the leading edge open to the cooling air
passage.
Trailing Edge Metal Loss
A burn through the convex trailing edge wall that causes a
shorter aerofoil chord length.
Coating Damage
Chips, spalling, flakes, blisters, peeling or oxidation that is
confined to the coating.
Engine InspectionlCheck

V2500 Borescope Practices Engine InspectionlCheck
Engine Inspectionlcheck
AMM ref. 72-00-00-200
This topic details the inspection procedures and
acceptance standards for the LP and HP compressors.
The combustion system, HP and LP turbine systems. The
HP nozzle guide vanes.
Inspection of these items requires the use of borescope
equipment and references to the Aircraft Maintenance
Manual, (AMM).
The accepthejectinformation can be found in AMM Ch 72-
00-00.
For inspection of other engine areas, reference must be
made to the Inspection/Check page block at the
appropriatechapter.
The following inspection/checkareas of the engine will be
discussed in this section;
0 LPC booster.
0 HP compressor.
0 Combustion system.
0 HPturbine.
0 LPturbine.
Note:
Engine rejection due to damage found being out of limits
must be reportedto the local IAE representative.
Any damage found not covered in the AMM must be
reported to the local IAE representative for further advice.
Inform your local IAE representative of any decrease in
the inspection intervals.
The limits quoted in the AMM are applicable on a continue
in service basis only.
The limits quoted in the manual are based on the condition
of the damage and the quantity of damage seen. The
limits are structured so as to allow safe continued
operation of the engine and hence the aircraft.
This is achieved by having reduced operation of the
engine, either in hours and/or cycles in between the
inspection intervals as the damage deteriorates.
Ifthe damage exceeds the ultimate limit then the engine is
scheduledfor removal as per AMM requirements.

0IAE InternatlonalAero EnginesAG 2000
IAE V2500 PROPULSIONUNIT

Q IAE internationalAero Engines AG 2000
V2500 Borescope Practices Engine Inspection/Check
LPC Booster InspectionlCheck
AMM ref. 72-00-00-200-017
LPC stage 1.5 (AI, A5, D5)
Guide tube IAE 2J12030 is used and entry is gained from
the L P C booster inlet guide vanes.
LPC stage 2.0 and 2.3 (A5, 05)
Removing twenty two fan blades: two fan exit guide vanes
and then removing a borescope blank to gain access.
LPC stage 2.5 (AI, A5, D5)
Access is gained by a fibrescope being passed into the
2.5 bleed outlet duct.
Each stage of the L P C booster has a number of blades
per disc. These are as follows:
Stage 1.5 52 blades 58 blades.
A I engines A5/D5engines
Stage 2.0
Stage 2.3
68 blades 78 blades.
88 blades.-------------
Stage 2.5 70 blades 72 blades.
Inspection Standards
The L P C booster can be inspected while the engine is on
wing.
The aircraft and engine must be in the prepared condition
as advised by the AMM prior to the inspection.
During the inspection the typical damage to look for is as
follows:
Examine for cracks.
0 Examine for nicks.
Examine for tears.
0 Examine for dents.
0 Examine for scratches.
Examine for tip damage, bends and rubs.
Any damage found can be assessed according to the
AMM recommendations.

2.5
LP COMPRESSOR BOOSTER INSPECTION

LPC Booster Inspection
The images below show how access is gained to the
stages of the LPC booster.
Stage 1.5 Compressor Inspection
One person is needed to locate the guide tube into the fan
stator vane, mark the fan for the start position and rotate
the fan for the inspection, take care to prevent any
damage to the Fan Exit Guide Vanes (FEGV) or bypass
duct structure.
One person is needed to insert the fibrescope and make
the inspection.
0 Install the guide tube (IAE 2J12030) at approximately
the 3 o’clock position viewed from the rear of the
engine.
0 Put the 6mm fibrescope into the guide tube until stage
1.5blades can be seen.
0 Mark a fan blade and the adjacent fan case using an
AMM approved marker for the start position.
Stage 2.0 and 2.3 Compressor Inspection (A51D5 Only)
Access to the 2.0/2.3 borescope plug can be gained by
carrying out the following tasks:
0 Remove the Inlet Cone (Spinner) and Fairing.
0 Remove the 22 Fan blades and Annulus fillers.
0 Remove the Outer Liner Panel and Splitter Fairing.
0 The 2 segments of Fan Exit Guide Vanes (FEGV) at
approximately 5 o’clock position viewed from the rear
of the engine.
0 Remove the borescope plug and key washer then
discard the key washer.
One person is needed to turn the LP system and mark the
start position for the inspection.
One person is needed to install the fibrescope and make
the inspection, for this position there is no guide tube.
0 Put the 6mm fibrescope into the access port until the
stage 2.0 and stage 2.3 blades can be seen.
0 Mark the LP system using an AMM approved marker
for the start position.
Stage 2.5 Compressor Inspection
One person is needed to turn the fan and mark the start
position for the inspection.
One person is needed to insert the fibrescope and make
the inspection.
0 Put the 6mm fibrescope into the outlet port of the LP
Compressor bleed valve until stage 2.5 blades can be
seen. At approximately the 6 o’clock position viewed
from the rear of the engine.
Mark a fan blade and the adjacent fan case using an
AMM approved marker for the start position.

Engine InspectionICheck
-
GUIDETUBES
P
1.5
7ON
AlIA51D5
STAGE
NSPECl
LPC BOOSTER INSPECTION

LPC Booster Inspection Blade Zones
Damage assessment of the L P C booster blades about the
aerofoil are identified into three separate zones. The
accept/reject limits for damage found in these zones may
vary between them.
The three zones are identified as:
0 ZoneA.
0 ZoneB.
0 ZoneC.
T o help the engineer to assess damage according to
which zone it is in the blade has been apportioned a
percentage of the total aerofoil surface.
The apportioning is as follows:
0 Zone A is 25% of the aerofoil from the root platform.
0 Zone B is 35% of the aerofoil above the zone A area.
0 Zone C is 40% of the aerofoil above the zone B area.
In order to assess damage in the zones according to AMM
recommendations the blade dimensions must be known.
The following is a table giving the blade dimensions at
each stage.
Height (2) True Width(1)
Stage inches mm inches mm
Blade 1.5-AI 3.51 89.1 1.71 43.5
Blade 1.5-A5/D5 3.78 96,O 1.52 38,6
Blade 2.0-A5/D5 3.08 78,3 1.20 30,5
Blade 2.3-A5/D5 2.95 74,9 1.10 28,O
Blade 2.5-AI 3.25 82.6 1.43 36.2
Blade 2.5-A5/D5 3.04 77,3 1.45 36,8
Note:
A I engines do not have a L P C booster stage 2.3.
AI engines stage 2.0 cannot be accessed for borescope
inspection.
The width is the true chord measurement.
The true width is measured along the chord line at middle
aerofoil height.

ZONE c
ZONE B
ZONE A
STAGE 1.5
A1/A5/D5
50% OF
AEROFOIL
2
1
ZONE C
A
t
STAGE 2.0 A5/D5
50% OF
AEROFOIL
A
2
ZONE B
ZONE A
4STAGE 2.3 A5/D5
2
7
1
-
ZONE C
ZONE B
50% OF
AEROFOIL
ZONE A
4STAGE 2.5 Al/A5/D5
LPC BLADE DIMENSIONSAND ZONES
2

HP Compressor InspectionlCheck
AMM ref. 72-00-00-200-016
A combination of using fibrescopes and rigid scopes is
recommended for the inspections.
0 Fibrescope size is 6.0mm.
0 Rigid scope sizes are 5.5mm and 8.0mm.
Each stage of the HP compressor has a number of blades
per disc. These are as follows:
Stage 3 31 blades 31 blades.
AI engines A5lD5 engines
AI engines A5/D5engines
Stage 8
Stage 9
Stage I O
Stage 11
The HP compressor can be inspected while the engine is
on wing. (For pre SB 72-0027 engines the AMM
recommends stage 3 and 12 only).
During the inspection the typical damage to look for is as
follows:
0 Examine for cracks.
0 Examine for nicks.
0 Examine for tears.
0 Examine for dents.
0 Examine for scratches.
0 Examine for tip damage, bends and rubs.
0 Examine for stage 1shingling, (snubber override).
0 Examine for ceramiccoating loss.
Any damage found can be assessed according to the
AMM recommendations.

I /
IMID HEIGHT SUPPORT VORTEX REDUCER
Rlhl
0
v)
HP COMPRESSOR STAGES 3 TO 12
8n

HP Compressor Inspection Blade Zones
Damage assessment of the HP compressor blades about
the aerofoil are identified into 3 separate zones for stages
4 through to 12.
Stage 3 is identified as 4 separate zones.
The acceptlreject limits for damage found in these zones
may vary between them.
The zones are identified as:
0 ZoneA.
0 ZoneB.
0 ZoneC.
HPC stage 3 has an extra zone known as Zone D.
dimensional size and/or percentage of the total aerofoil
surface.
HPC Stage 3
The HPC stage 3 blade is apportioned into four zones.
They are as follows:
Zone A is the remainder above zone B and to the blade
tip.
Zone B is the area above and below the mid span support.
(Zone B is 6.5mm above and 6.5mm below the mid span
Zone C is the remainder from the top of zone D to the mid
span support.
Zone D is 13mm (0.5in) from the blade root up.
HPC Stage 4 to 6
Zone A is 30% of the blade aerofoil surface from the blade
tip.
Zone B is the remainder of the aerofoil surface between
zones A and C.
Zone C is 13mm (0511) from the blade root and upwards.
HPC Stage 7 to 12
Zone A is 30% of the blade aerofoil surface from the blade
tip.
Zone B is the remainder of the aerofoil surface between
zones A and C.
Zone C is 6,4mm (0.25in) from the blade root and
upwards.
support).

0IAE InternalionalAero Engines AG 2000
HEIGHT AT MIDSPAN IS
6,5mm (0.25in) ABOVE AND
BELOW THE CENTER LINE
HPC STAGE 3 /
HPC STAGE 7 TO 12
I I
HEIGHT IS 13mm (0.5in) HEIGHT IS 6,4mm (0.25in)LHElGHT IS 13mm (0.5in)
HP COMPRESSOR BLADE ZONES

HP Compressor Inspection Blade Dimensions
In order to assess damage in the zones according to AMM
recommendationsthe blade dimensions must be known.
The following is a table giving the blade dimensions at
each stage.
Dim X Dim Y Dim Z
Stage
3
4
5
6
7
8
9
10
11
12
inches
5.04
3.66
2.62
1.9
1.45
1.I4
0.94
0.86
0.83
0.84
mm
128,2
93,l
66,5
48,3
36.9
29,O
23,8
22,o
21,o
21,2
inches
2.39
2.20
1.57
1.15
0.92
0.91
0.86
0.85
0.89
0.86
mm
60,7
55,8
39,8
29,3
23,4
23,2
21,8
21,6
22,5
21,8
inches
2.54
I.73
1.32
0.97
0.74
0.59
0.48
0.45
0.43
0.44
mm
64,6
44,O
33,6
24,6
18,8
14,9
12,3
11,4
10,9
11,l
Note:
These dimensions are for reference purposes only.
The chord dimension changes from the blade root to the
bladetip.
The chord datum (dimension Y) is at dimension Z
measured from the blade root.
Dimension X is measured at the mid chord position.
Engine InspectionKheck

. .
LEADINGEDGE
f
HP COMPRESSOR BLADE DIMENSIONS

HP Compressor Inspection
HPC Blade Tip Rubbing
The HPC has exhibited blade tip rubbing. If you suspect
Ceramic Coating Detachment
Ceramic coating detachment is
that blade tip rubbing has occurred then consult the AMM
for further advice.
The AMM outlines limits for blade tip rub at stages 3 to 8
that are different to blade tip rub limits for stages 9 to 12.
Damper Wire Detachment
HPC stages 7 rear and 8 front have damper wires installed
to the blades. If you suspect that the damper wires have
become detached consult the AMM for further advice.
The damper wires of pre modification SB 72-0289 HPC's
have exhibited damper wire detachment. This can lead to
secondary engine damage of the HPC.
SIL 106;
SB 72-0289 introduced a fix for the existing style of HPC
disc that modifies the existing wires with rounded ends.
SB 72-0304 introduces a new type of wire. The wires are L
shaped for further improved reliability.
SB 72-0300 introduces a new design of disc that has
changed and improved the reliability of the wires.
Use borescope access ports C, D and E for examination
for missing wires.
acceptable as long as the
limits in the AMM are adhered to.
Ceramic coating loss from the compressors has been
linked to the events that have caused the erroneous input
pressure signal of Pb to the EEC.
This is where fine particles of ceramic material have
entered the Pb pressure tube and caused contamination
of the tube and thus affecting the pressure signal.
Cracks, Nicks,Tears and Dents
The limits for the above title in the AMM varies according
to the stage of compressor that is being inspected.
They are as follows:
Stage 3.
0 Stage4 to6.
0 Stage 7 to 12.
There are accept and reject standards for the above
mentioned HPC stages.

..,.-.. .. .......
0 IAE International Aero EnginesAG 2000
MATERIAL LOSS
STATOR BLADE
MID CHORD TIP RUB

CERAMICCOATED
ABRAIDABLE STATOR PATH
2
st
n HPC TIP RUB AND DAMPER WIRE DETACHMENT

Combustion Chamber and HPT Stage INGV's
The combustion chamber and the components that are
operating along with it are subjected to the highest
temperatures and pressures within the engine.
Therefore the inspection requirements are on a more
regular basis according to the MPD schedule
requirement s.
A typical combustor and turbine inspection is normally
scheduled in accordance to the maintenance practices
document (MPD).
The areas that are inspected within the combustion
system are as follows:
Borescopeaccess
Access for combustion chamber inspection is as follows:
The igniter plug port positions of IP1and IP2.
Combustion chamber ports B1through to B5 (Pre SB 72-
Combustion chamber ports B1through to B6 (Post SB 72-
Note:
After engine shut down a hot engine can cause damage to
borescope inspection equipment. It is advised that a time
interval of 2 to 3 hours be allowed to pass prior to
0221).
0221).
commencingengine inspection.
Inner liner shell.
Outer liner shell.
Inner burner liner segment.
Outer burner liner segment.
Bulkhead segment.
Bulkhead deflector.
Fuel spray nozzles (FSN's).
HPT stage 1 NGV's.
The inner and outer burner liner segment rows are
numbered 1to 5 starting at the FSN's.
The AMM has printed forms for the recording of damage
found.

COMBUSTION CHAMBER BORESCOPE VIEWS

Inspection of the combustion system and HPT stage 1
NGV's are as follows:
Fuel spray nozzles
The inspection requires that the FSN deflector end cap
and the critical areas of the nozzle deflector adjacent to
the cap be inspected for;
0 Cracks.
0 Distortion.
Erosion.
0 Broken.
0 Burned.
Combustion System
The combustion system should be examined for the
following:
Burns.
Burn holes.
Edge burnback.
Radial burnback.
HPT 1'' Stage NGV's
The HPT 1'' stage NGV's should be inspected for the
following:
0 Burns.
Burn through.
0 Trailing edge burns.
0 Lift up.
0 Coatingdamage.
0 Connectedloop crack.
Note:
Damage assessment can be greatly enhanced by using
parts within the combustion system that are of a known
dimension.
Use parts that are of a known dimension that are nearest
to the damage.
Special Inspectionto SB 72-0349
This is an A1 and early A5 series of engines inspection
requirement. The inspection is as follows:
HPT stage 1 rotor metering plugs for heat distress/oil
wetness.

0IAE International Aero EnginesAG 2000
COMBUSTOROUTER
LINER SHELL
I
OUTER LINER
SEGMENT
' COMBUSTION CHAMBER AREAS OF INSPECTION

Q IAE internationalAero EnginesAG 2000
V2500 BorescopePractices Engine Inspection/Check
Fuel Spray Nozzle (FSN) Inspection
The FSN inspection requirements in the AMM ask for
inspection of the end caps.
Inspection of the bulkhead segments and the bulkhead
deflectors are covered in the AMM giving limits for accept
and reject standards.
The following are the inspection requirements for the FSN,
bulkhead deflectors and bulkhead segments.
Fuel Spray Nozzle
There are 20 FSN's that require inspection for the
following:
0 End caps.
End cap damage can be accepted as long as the limits in
the AMM are adhered to.
Bulkhead Segments
There are 20 bulkhead segments that require inspection
for the following:
0 Tight cracks.
0 Burns.
Tight cracks and burns damage can be accepted as long
as the limits in the AMM are adhered to.
Bulkhead Deflector
The bulkhead deflectors are on the bulkhead segment
main body. These require inspectionfor the following:
0 Cracks.
0 Burns.
Cracks and burns damage can be accepted as long as the
limits in the AMM are adhered to.
Note:
The limits quoted in the manual are based on the condition
of the damage and the quantity of damage seen. The
limits are structured so as to allow safe continued
operation of the engine and hence the aircraft.
This is achived by having reduced operation of the engine,
either in hours and/or cycles in between the inspection
intervals as the damage deteriorates.
If the damage exceeds the ultimate limit then the engine is
scheduled for removal as per AMM requirements.
initial issue Page4-23

TYPICALVIEWS OF END CAP DAMAGE
LOSS RADIAL CRACKS
CIRCUMFERENTIAL
CRACKS
CAP
FUEL SPRAY
BULKHEAD SEGMENT
FUEL SPRAY NOZZLE
FUEL SPRAY NOZZLE AND BULKHEAD SEGMENT
Page 4-24Initial issue

Combustor Inspection
The inspection requirements for the combustor in the Note:
AMM ask for the inspection ofthe following:
0 Inner burner segment.
0 Outer burner segment.
Inner liner shell.
0 Outer liner shell.
Inspection of the combustor is covered in the AMM giving
limits for accept and reject standards.
The following are the inspection requirements for the
combustor:
Inner and Outer Burner Segments
0 Combustion holes.
0 Segment cracks.
0 Segment edge burnback.
0 Surface burns.
0 Burn holes.
The following damage listed above can be accepted as
long as the AMM limits are adheredto.
Inner and Outer Liner Shell
The inspection criteria are as for the inner and outer
burner segments. The limits for accept and reject
standards are found with the segments limits. The AMM
states where applicable the possible damage for the
Iiners.
Initial issue
If a piece of combustion chamber breaks off in between
inspection times then a borescope inspection of the HPT
is recommended.
If you notice spalled ceramic coating on the combustor
segments, bulkhead segments and bulkhead deflectors
this is acceptable without a decreased inspection interval.
Page4-25

Initial issue
C.C. IBL & OBL SEGMENT DIMENSIONS _I
Page 4-26

Stage 1 HPNGV Inspection
The inspection standards for the stage 1 HPT NGV’s Note:
All ceramic coating decrease is acceptable without arequire the following areas to be inspected:
0 Concave aerofoil surface. decrease borescopeinspection interval.
0 Convex aerofoil surface. Extended operation with a large quantity of ceramic
coating not there can possibly cause the vane to be out of
0 Inner platform.
limits.
0 Outer platform.
0 Aerofoil leading edge.
0 Aerofoil trailing edge.
Inspection of the HPT NGV’s is covered in the AMM giving
limits for accept and reject standards.
The following are the inspection requirements for the HPT
NGV’s:
0 Cracks.
0 Burns.
Burn through.
0 Trailing edge burns.
0 Lift up.
0 Coating damage.
0 Connected loop crack.

I
OD PLATFORM
LEADING EDGE
1/2 DISTANCE
DISTANCE I
1
Initial issue
2.100in
(53,34mm) (43,18mm)
PLATFORM LEADING EDGE
COOLING
AIR HOLES
HPT STAGE 1VANE DIMENSIONS A
Page 4-28

HPT Stage 1 Rotor Metering Plug Inspection
SB 72-0349 covers the requirement for the inspection of
the rotor metering plug. This SB affects the following
engines:
A I Engines before s/nV0362.
A5 Engines before s/n V10080.
During certain windmilling conditions it is possible for
engine oil to flood the n0.4 bearing compartment and
collect in the HPT near the stage 1 rotating air seal.
Subsequent engine operation may lead to turbine distress.
Inspection requirements
The equipment required to do this procedure is;
0 A 3mm (maximum diameter) flexi scope.
0 Guide tube (tool no.NDIP-988-GT).
In addition to the above special equipment tools are
required to remove the 1OIh stage nut and additional
borescope equipment is required to support the flexi
scope.
The entry point of the engine is located on the diffuser
casing left hand side. The tube connection which requires
removing for this task is the HPC stage 10 supplementary
air coming from the make up valve.
The guide tube allows the flexi scope to reach up to the
n0.4 bearing support cooling hole.
From this point and on the flexi scope will have to be
worked into position to view the metering plugs.

PATH OFTHE FLEX1SCOPE
I
No4 BEARING
SUPPORT,
COOLING HOLE
No 4 BEARING
m
w
loTHSTAGE MAKE U
STAGE 10 MAKE UP AIR
VALVE OFF TAKE TUBES
I-
ENTRY POINT FOR
INSPECTION
METERING PLUGS
LOCAT1ON
INSPECTION OF THE HPT STAGE IMETERING PLUGS

High Pressure Turbine (HPT) Stages 1 and 2
HPT Stage 1 ref., AMM 72-00-00-200-014
HPT Stage 2 ref., AMM 72-00-00-200-015
The HPT is a 2 stage turbine system designed to drive the
HPC.
Borescope access to the HPT is possible for both stages.
The recommended inspection equipment to effect a
borescope inspection of the HPT is as follows:
0 Rigid borescope of 8mm to give a general inspection
for damage.
Flexible borescope of 6mm for sufficient detail
inspection.
The borescope access ports are located on both sides of
the core engine. They are identified as follows:
HPT Stage 1 Leading Edge
0 IP1and/or IP2.
0 B1through to B5 (Pre SB 72-0221).
0 BI through to B6 (Post SB 72-0221).
HPT Stage 1Trailing Edge, Stage 2 Leading Edge
0 T1/2L which is on the left hand side.
0 T1/2Rwhich is on the right hand side.
HPT Stage 2 Trailing Edge
0 T2/3L that is the left hand side.
0 T2/3R that is the right hand side.
Initial issue
Each stage of the HP turbine has a number of blades per
disc. These are as follows:
0 HPT stage Ihas 64 blades.
0 HPT stage 2 has 72 blades.
Inspect the HPTfor the following:
0 Erosion.
0 Cracks.
0 Nicks.
0 Dents.
0 Burns.
0 Leading edge burn through.
0 Trailing edge metal loss.
0 Coating damage.
Accept and reject standards can be found in the AMM for
damage found in the engine.
Page4-31

FUEL SPRAY NOZZLE
HPT 1 LIE-FLEX1
ACCESS THROUGH
IGNITER PORT
HPT 1 UE-RIGID
L/E=LEADINGEDGE ACCESS THROUGH
T/E=TRAILINGEDGE COMBUSTORCASE
00 HPT2 LIE
8
m
v)
HPT STAGES IAND 2 BORESCOPE VIEWS
2W
0

HP Turbine Stage 1 Blade Zones
Damage assessment of the HP turbine blades about the
aerofoil are identified into separate zones.
The acceptheject limits for damage found in these zones
may vary betweenthem.
The zones are identified as:
0 ZoneA.
0 ZoneB.
0 ZoneC.
0 ZoneD.
surface.
Zone A
Zone A is 25% of the blade area from the root upwards.
Zone A is the leading edge moving inwards by 0.125ins
(3,175mm).
Zone A is the trailing edge moving inwards by 0.2ins
(5,08mm).
Zone B
Zone B is 50% of the blade area above the zone A area
and within the zone A leading and trailing edge areas.
Zone C
Zone C is the remaining area of the blade above the zone
B area and within the zone A trailing and leading edge
areas.
Zone D
Zone D is the tip area on the convex side only for A I
engines.
Blade Dimensions
The HPturbine stage 1blade overall radial length is:
0 100% equals 1.80in (45,72mm).

0.187 in
(4,75 mm)
t
NOTE: 100% SPAN EQUALS
1.70in (43,18mm)
DEV250239
ZONE D
1
-
100%
STAGE 1 HPT BLADECONVEX AIRFOIL VIEW STAGE 1HPT BLADE CONCAVEAIRFOIL VIEW
HPT STAGE I BLADE ZONES

0IAE InternalionatAero Englnes AG 2000
HP Turbine Stage 1 Blade Inspection
The AMM gives flexibility with the assessment of the HP
turbine. As the damage worsens then the inspection
interval will decrease andlor the engine hourskycles
decrease. This will be the acceptable practice until the
ultimate limit has been reached which will then require the
engine to be removed.
Damage assessment limits may vary from one zone to
another zone. Where this is applicable the AMM will
advise.
Inspect the HPturbine blades for the following conditions:
Cracks, Nicks, Burns and Dents on the Airfoil Surface
Generally cracks, nicks, burns and dents are acceptable
as long as the AMM limits and procedures are adhered to.
Cracks and Burns on the Platform of Stage 1 HPT
Blade
Generally cracks and burns on the platform of the stage 1
HPT blade are acceptable as long as the AMM limits and
procedures are adhered to.
Stage 1 HPT Duct Segment for Burn Holes
Generally burn holes are acceptable as long as the AMM
limits and procedures are adhered to.
Note:
When you do a borescope inspection of the stage 1HPT
blades some areas of the stage 1HPT duct segments can
be seen.
Engine InspectiorKheck

V2500 BorescopePractices EngineInspectionlCheck
ROOT PLATFORMCRACK
0 STAGE 2 NGVs BLADE PLATFORM
80
In
8
HPT STAGE I ROTOR BLADE AND DUCT SEGMENT

HP Turbine Stage 2 Blade Zones
Damage assessment of the HP turbine blades about the
aerofoil are identified into 4 separate areas.
The acceptheject limits for damage found in these areas
The areas in question are called zones. The zones are
identifiedas:
0 ZoneA.
ZoneB.
0 ZoneC.
0 ZoneD.
surface.
Zone A
Zone A is 50% of the aerofoil length from the root
upwards. Note that the area of A starts at 0.250ins
(6,35mm) from the root even though the overall length of A
is taken from the root.
Zone A is O.5ins (12,7mm) in from the leading edge.
Zone B
Zone B is 25% of the aerofoil length above the area of
zone A and 0.5ins (12,7mm) in from the leading edge.
Zone C
Zone C is 25% of the aerofoil length from the tip
downwards.
Zone D
Zone D is 75% of the aerofoil length from the root upwards
which does not encroach into zones A andlor B.
Blade Dimensions
The HP turbine stage 2 blade overall radial length is:
0 100% equals 2.50in (63,50mm).
initial issue Page4-37

0IAE lnternatlonal Aero Engines AG 2000
AXIAL DIRECTION
Initialissue
75%
50%
HP TURBINE STAGE 2 BLADE ZONES
Page4-38

'
HP Turbine Stage 2 Blade Inspection
The AMM gives flexibility with the assessment of the HP
turbine. A s the damage worsens then the inspection
interval will decrease and/or the engine hourslcycles
decrease. This will be the acceptable practice until the
ultimate limit has been reached which will then require the
engine to be removed.
Damage assessment limits may vary from one zone to
another zone. Where this is applicable the AMM will
advise.
Inspect the HPturbine blades for the following conditions:
Cracks, Nicks, Erosion, Burns, holes and Dents on the
Airfoil Surface
Generally cracks, nicks, erosion, burns, holes and dents
are acceptable as long as the AMM limits and procedures
are adhered to.
Cracks and Burns on the Platform of Stage 2 HPT
Blade
Generally cracks and burns on the platform of the stage 2
HPT blade are acceptable as long as the AMM limits and
procedures are adhered to.
Stage 1 HPT Duct Segment for Burn Holes
Generally burn holes are acceptable as long as the AMM
limits and procedures are adhered to.
Note:
When you do a borescope inspection of the stage 2 HPT
blades some areas of the stage 2 HPT duct segments can
be seen.

BLADE TIP CRACKS
LEADING EDGE
AEROFOIL BURNS
TRAILING EDGE
UE NICK WITH A CRACK
ROOT PLATFO AEROFOIL BURNS
WITH CRACKS
HPT STAGE 2
DUCT SEGMENT
HPTSTAGE 2 NGVs
hl
e
In
8
QE

BLADE ROOT PLATFORM
HPT STAGE 2 BLADE AND DUCT SEGMENT

LP Turbine Stage 3 InspectionlCheck
AMM ref. 72-00-00-200-012
The LPT stage 3 rotor blade leading edge is accessible for
inspection. There is a borescope port provided.
The LPT stage 7 trailing edge is accessible for inspection.
This is viewed by looking down the exhaust nozzle.
For LPT stage 3 borescope access port T112L and T1/2R
can be used.
The recommended inspection equipment to effect a
borescope inspection of the LPT stage 3 is as follows:
0 Rigid borescope of 8mm to effect an inspection for
damage.
0 Flexible borescope of 6mm to effect an inspection for
damage.
There is also a requirement for the borescope access plug
to be inspected.
It must be noted that the spacing washer must be kept
with the borescope plug and not allowed to mix with other
spacing washers.
The inspection standards for the LPT stage 3 are as
follows:
0 Cracks.
0 Nicks.
0 Tears.
0 Dents.
Initialissue
0 Bends.
0 Blistering or loss of coating.
0 Burning or oxidation.
0 Build up of deposits.
0 Sulphidation.
0 Flame plate extrusion.
Rotor Blades
There are 97 rotor blades about the disc for LPT stage 3.
There are 89 rotor blades about the disc for LPT stage 7.
Page4-41

LPT STAGE 3 STATOR VANE I//I
LPT
BORESCOPE PROBE
STAGE 3 ROTOR BLADE BORESCOPE
LPT STAGE 3
ROTOR BLADE
3
VIEW

0 IAE lnternalional Aero Engines AG 2000
LP Turbine Stage 3 Blade Zones
Damage assessment of the LP turbine blades about the
aerofoil are identified into 3 separate areas.
The acceptlreject limits for damage found in these areas
The areas in question are called zones. The zones are
identified as:
0 ZoneAl.
0 ZoneA2.
ZoneA3.
surface area.
Zone A I
Zone A I is O.lins (2,5mm) from the blade shroud moving
towards the root.
Zone A2
Zone A2 is the majority of the blade aerofoil surface.
This is the area below the zone A I limit and above the
zone A3 limit.
Zone A3
Zone A3 i s 0.12Oins (3,Omm) from the blade root moving
upwards.

V2500 bsi issue 01

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