V2500 bsi issue 01

I--

V2500 ABBREVIATION ACAC

ACC

ACOC

A I D R S

Alt

A P U

AMM

BDC

BMC

BSBV

CFDIU

C F D S

C L

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 Management Computer

Electronic Centralised Aircraft Monitoring

Environmental Control System

Electronic Engine Control

EHSV

EIU

EIS

E V M S

EVMU

EPR

E T O P S

FADEC

FAV

FCOC

F C U

FDRV

FSN

FMGC FMV

F M U

F O B

F W C

H C U

H IV

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

Ful l 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 Management and Guidance Computer Fuel Metering Valve

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

I DG

IAE

I DG

IFSD

IGV

I bs.

LE

LGCl U

LGCU

LH

LP

LPC

LPCBV

LPSOV

LPT

LRU

LT

LVDT

MCD

MCDU

High Pressure Compressor

High Pressure Turbine

High Pressure Regulating Valve

High Tension (ignition lead)

Integrated Drive Generator

International Aero 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 Pressure Compressor

Low Pressure Compressor Bleed Valve

Low Pressure Shut off Valve

Low Pressure Turbine

Line Replaceable Unit

Low Tension

Linear Voltage Differential Transformer

Magnetic Chip Detector

Multipurpose Control and Display Unit

MCLB

MCT

Mn

MS

NAC

NGV

NRV

N I

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-Return Valve

Low Pressure system speed

High Pressure system speed

Outside Air Temperature

Outlet Guide Vane

Open

Over Pressure Valve

Overs peed

Pressure Ambient

Burner Pressure

Pressure Regulating Shut Off Valve

Pressure Regulating Valve

Pounds Per Square Inch

Pounds Per Square Inch Differential

Permanent Magnet Alternator

Pressure of the fan inlet

Pressure of the LP compressor outlet

Pressure of the HP compressor outlet

P4.9

QAD

SAT

SEC

STS

TA1

TAT

TAP

TCT

TDC

TE

TEC

TFU

TRA

TLA

TLT

TM

TO

Pressure of the LP turbine outlet

Quick AttachlDetach

Static Air Temperature

Spoiler Elevator Computer

Status

Thermal Anti Ice

Throttle Angle Transducer

Transient Acoustic Propagation

Temperature Controlling Thermostat

Top Dead Centre

Trailing Edge

Turbine Exhaust Case

Transient Fuel Unit

Throttle Resolver Angle

Throttle Lever Angle

Temperature Limiting Thermostat

Torque Motor

Ta ke-off

.. .. . .. .,. .

INTERNATIONAL AERO ENGINES AG V2500 BORESCOPE PRACTICES CONTENTS

Section 1

Section 2 Borescope Requirements.

Section 3 Engine Maintenance Practices.

Section 4 Engine Inspection/Check.

Section 5

Section 6 Typical Examples

Introduction and Engine Mechanical Arrangement.

Part ConditionlClassification of Damage.

INTRODUCTION AND ENGINE MECHANICAL ARRANGEMENT

0 IAE International Aero Engines AG 2000 IAE V2500 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. T h e 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.

T h e 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:

Introduction and Mechanical Arrangement

Initial issue Page 1-1

8 IAE International Aero Engines AG 2000 IAE V2500 Borescope Practices Introduction and Mechanical Arrangement

IAE International Aero Engines AG

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 are:

0 Rol ls 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 I A E 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 Rol ls Royce plc--hig h 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: Rol l s Royce have introduced the wide chord fan to the V2500 engine family.

The senior partners Rol ls 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.



IAE V2500 EnginelAirframe Applications

T h e V2500 engine has been designated the V because International Aero Engines (IAE) were originally a five nation consortium. T h e V being the Roman numeral for five.

T h e 2500 numbering indicates the first engine type to be released into production rated at 250001bs of thrust.

F o r easy identification for the present and future variants of the V2500 A5/D5, I A E introduced the following engine designation system.

All engines will retain the V2500 numbering a s a generic name.

T h e first three characters of the full designation are V25. T h i s will identify all the engines in the family.

T h e next two figures indicate the engines rated sea level takeoff thrust.

T h e following letter shows the aircraft manufacturer.

T h e last figure represents the mechanical standard of the engine.


T h e designation V2500-D collectively describes all applications for the Boeing McDonnell Douglas aircraft.

T h e V2500-A collectively describes all the applications for the Airbus lndustrie aircraft.

This is irrespective of engine thrust rating.

T h e number given after the alpha indicates the mechanical standard of the engine. F o r example;

T h e V2500 Alengine is exempt from these idents a s it was certified with only one thrust rating.

V2528-A5.

This system will provide a clear designation of a particular engine as well a s a simple way of grouping by name engines with similar characteristics.


IAE V2500 Borescope Practices 0 IAE International Aero Engines AG 2000


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 THRUST 250001bs

V2524 AS-EIS1 997 AIRBUS A319 THRUST 240001bs

BOEING MD90 THRUST 250001bs

0 V2525 D5-EIS1995

V2522 ABEIS AIRBUS A319 THRUST 220001bs

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/AIRFRAME APPLICATIONS t; 0




Introduction to the Propulsion System T h e V2500 family of engines share a common design feature for the propulsion system.

T h e complete propulsion system comprises the engine and the nacelle. T h e major components of the nacelle are:

0 T h e intake cowl.

0 T h e fan cowl doors.

0 Hinged C ducts with integral thrust reverser units.

0 Common nozzle assembly.

Intake Cowl

T h e intake cowl allows the smooth intake of air to the engine while providing an aerodynamic exterior to reduce engine drag.

T h e intake cowl contains the minimum of accessories. T h e 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.

T h e fan cowl doors are hinged to the aircraft strut in four positions and secured in the closed position by four latches.

T h e 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

T h e 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)

T h e 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 Introduction and Mechanical Arrangement

IAE V2500 PROPULSION UNIT Initial issue Page 1-6



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 Mechanical Arrangement 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 handling characteristics.

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.



COMBUSTOR HP COMPRESSOR I

BOOSTER

LP COMPRESSOR

AIRCRAFT STRUT

A51D5

0 0

In ft 2 W 0

ACCESSORY GEARBOX

PROPULSION SYSTEM OUTLINE Initial issue Page 1-8

Q IAE International Aero Engines AG 2000 IAE V2500 Borescope Practices Introduction and Mechanical Arrangement

Engine

Active clearance control (ACC) turbine

Active clearance control (ACC) is used on both the LP and HP turbine casings.

T h i s 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 LP turbine 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 Automatic engine starting.

0 Compressor airflow control system.

0 Heat Management system.

0 10th stage make up air system.

0 Thrust reverser.

The E E C also provides protection for:

0 N1 overspeed.

0 N2 overspeed.

Engine surge.


IAE V2500 Borescope Practices

c

(0 IAE lnternatlonal Aero Engines AG 2OOO Introduction and Mechanical Arrangement

1 W

cl

Initial issue

V2500 ENGINE CUT AWAY Page 1-10

@ IAE International Aero Engines AG 2000 IAE V2500 Borescope Practices

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 - IntakelEngine inlet 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. Propulsion Unit Data

T/O thrust (SL static) Flat rated temperature Total airflow By pass ratio Overall pressure ratio Fan diameter Propulsion unit length Engine overall length Propulsion unit weight Bare engine weight

250001bs (1 I 1205kn) 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 (331 1 kgs) 49421 bs (2242kgs)

Page 1-1 1 Initial issue



A I

A51D5

HP COMPRESSOR STAGES I l l I 1.5 2 2.3 2.5

LPC BOOSTER A5/D5 U, / 8 LPCSTAGEI 5 t P ENGINE STATION AND STAGE NUMBERING


0 IAE international Aero Engines AG 2000 IAE V2500 Borescope Practices

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 Delivery Temperature).

P3 (or Pb) Pressure at 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.

Tern perat u re

Engine Gas Temperature (EGT) is T4.9.

Trend Monitoring

Trend Monitoring uses signals of P12.5, T2 and T3.

Note: Stations 4 81 4.5 are not sensed.





LPTPlSA

SLTO PRESSURE - TEMPERATURE MAP (A1 engine flgs for 25,0001bs thrust)

126.0 in-

DIMENSIONS

ENGINE DIMENSIONS AND PRESSURE TEMPERATURE MAP



Module no. 31

0 IAE International Aero Engines AG 2000 Introduction and Mechanical Arrangement

Module Fan

Introduction-Module Breakdown of the Engine The engine is of a modular construction and consists of Modular construction gives the following benefits:

40 50

HP System LP Turbine

I 3 2 I Intermediate Module

r 60 I External Gearbox

0 Lower overall maintenance costs.

0 Maximum life achieved from 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 independently balanced.

Note: Module numbers refer to the A T A section number of the Chapter/Section/Subject reference in the manuals. The HP system 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 - HP Turbine

Initial issue Page 1-1 5


r 31 -FAN 32-INTERMEDIATE

60-EXTERNAL GEARBOX

ENGINE MODULES initial issue Page 1-1 6



Module 31-Low Pressure Compressor (LPC)

Purpose

T h e low pressure compressor is designed to move a large volume of air rearwards, which constitutes to the majority of the engines thrust.

Location

T h e low pressure compressor is attached to the L P C stubshaft via a curvic coupling. T h e L P C stubshaft is attached to the LPT shaft.

Description

T h e low pressure compressor is the complete fan assembly and comprises:

0 22 Hollow fan blades.

0 22 annulus fillers.

0 T h e fan disc. 0 T h e front and rear blade retaining rings.

T h e blades are retained in the disc radially by the dovetail root with the front and rear blade retaining rings providing axial retention. Blade removal/replacement is achieved by removing the front blade retaining ring and the annulus fil lers 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.

T h e nose cone and fairing smooth the airflow into the fan but do not form part of the module 31 they are non- modular parts.


<

IAE V2500 Borescope Practices 0 IAE Internatlonal Aero Engines AG 2000


REAR BLADE

FRONT BLADE

NOSE CONE

LP COMPRESSOR (FAN) L



Module 32 Intermediate Module

Purpose

T h e intermediate module provides an airflow link between the L P C booster and HP system. T h e intermediate module houses the front bearing compartment and the drive between gearbox and the core engine.

T h e fan casing provides a path for the fan to operate in and also provides protection against a singular fan blade failure.

Location

T h e intermediate module is the main mounting location for the other engine modules.

Description

T h e Intermediate Module comprises of:

0 Fancase.

0 Fanduct.

0 Fan outlet guide vanes.

0 LP compressor booster (A5 - 4 stage)(Al - 3 stage).

0 LP compressor 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 LP stub shaft.


0 Inner support struts.

0 Outer support struts.

0 Vee groove locations for the inner and outer barrels of the 'C' d u cts .



Initial issue' '

INTERMEDIATE MODULE FORWARD VIEW 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

E 0

Page 1-22


Module 40 HP Compressor

Purpose

T h e H P C is designed to provide the combustor with high pressure compressed air.

Location

T h e H P C is located between the L P C booster and the com bustor.

Description

T h e HP compressor assembly is a 10 stage axial flow compressor.

It has a rotor assembly and stator case. T h e stator case gives support for the rotor assembly.

T h e compressor stages are numbered from the front starting at the L P C booster inlet. This means that the H P C 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.

1 stageVIGV.

3 stages VSV.

A5 standard has:

T h e rotor assembly has five sub-assemblies:

0 Stages 3 to 8 HP compressor disks.

0 A vortex reducer ring.

0 Stages 9 to 12 HP compressor disks.

0 T h e HP compressor shaft.

0 T h e HP compressor rotating air seal.

T h e five sub-assemblies are bolted together to make the rotor.

The compressor blades in stages 3 to 5 are attached to the compressor d i sks in axial dovetail slots and secured by lockplates.

T h e 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:

T h e HP compressor front case.

T h e HP compressor rear case.

All stages of the H P C can be accessed by borescope inspection. T h e majority of the blanks, ports and cover plates are located on the right hand side of the engine.




r: 8 2 W n

Initial issue

'ER

CASE

HP COMPRESSOR ASSEMBLY Page 1-24



Compressor Drums (Rotor)

The rotor assembly is constructed in two parts:

0 The stages 3 to 8 drum.

0 The stages 9 to 12 drum.

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.



0 IAE International Aero Engines AG 2000 introduction and Mechanical Arrangement

initial issue HP COMPRESSOR DRUMS

Page 1-26



Compressor Blades

T h e high pressure compressor is made up of ten stages of compressor type blades.

0 T h e first stage is identified as stage 3.

0 T h e last stage is identified a s stage 12.

T h e compressor blades in stages 3 to 5 are attached to the discs in axial dovetail slots and secured by lock plates.

Rubber str ips bonded to the underside of the platform seal gaps between the blades.

T h e 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 dLocKNuT STAGES 4 TO

5 BLADES

\ \ \ I

LOCK SCREW

HPC STAGE 3 ’

&\\ LOCK BLADES

AXIAL DOVETAIL SLOT

CIRCUMFERENTIAL DOVETAIL SLOTS

HP COMPRESSOR BLADES Initial issue Page 1-28

Q IAE International Aero Engines AG 2000 Introduction and Mechanical Arrangement IAE V2500 Borescope Practices

Module 40 HP System

Com bustor

Purpose

The combustor is designed to mix the compressed air with fuel and ignite the mix. The hot gasses are then expanded to drive the turbines.

Location

The combustor is located between the HPC and HPT.

Description

The combustion system 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 10 F4

E1 t; 0

Initial issue

OUTER COMBUSTION i LINER LINER SEGMENTS

COMBUSTOR SYSTEM

JSTION

Page 1-30


Module 40 HP System

Corn bustor

The drawing below shows the arrangement of:

0 The diffuser casing.

The inner and outer combustion liners.

The HPT stage 1 NGV’s.

0 The T O B I (Tangential out Board Injector).

Also shown are the No 4 bearing support assembly and the primary parts of the stage 1 turbine nozzle assembly (HP NGV)

The following components make up the inner liner assembly of the combustion chamber:

0 The Stage I HPT Vane Cluster Assemblies.

0 The Stage 1 HPT Cooling Duct Assembly.

0 The Combustion Chamber Inner Liner.

0 The stage 1 turbine 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 1 HPT cooling duct assembly with bolts. The hollow vane airfoils have internal baffles and cooling holes.




HPC OUT1 VANES

COMBUSTION CHAMBER OUTER LINER

COMBUSTION I CHAMBER INNER LINER FUEL SPRAY

NOZZLE

DIFFUSER CASE // ASSEMBLY

No 4 BEARING / SUPPORT ASEMBLY

COMBUSTOR CROSS SECTION



Module 40 HP System

HP Turbine (HPT)

0 IAE International Aero Engines AG 2000

Purpose

T h e 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 H P C a s a result.

Location

T h e HPT is located between the combustor unit and the LPT stage 3.

Description

T h e primary parts of the HP turbine rotor and stator assembly are:

0 T h e HP Turbine Rotor Assemblies (Stage 1 and Stage

0 T h e HP Turbine Case and Vane Assembly.

T h e HP turbine rotor assemblies are two stages of turbine hubs with single-crystal, nickel-alloy blades with high strength and resistance to creep. T h e 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).


T h e primary parts of the stage 1 rotor assembly are:

Stage 1 turbine hub.

0 Inner and outer HPT air seals.

0 64 Blades.

0 Rear HPT air seal.

T h e 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 1 and 2.



IAE Introduction and Mechanical Arrangement V2500 Borescope Practices

HPC loTH STAGE AIR FOR

HPT STAGE 2 HPT STAGE 1

HP NGV 2

HPT COOLING AIR FEED TO STAGE 1

HPC 1 OTH STAGE AND STAGE 2 HPT DISCHARGE AIR TO HPT AIRSEAL

Initial issue HP TURBINE ASSEMBLY

Page 1-34


Module 50 LP Turbine

Purpose

Q IAE International Aero Engines AG 2000 Introduction and Mechanical Arrangement

T h e low pressure turbine (LPT) i s designed to maintain the rotational momentum of the L P C system.

Location

T h e LP system module is attached to the HP system module and is linked to the L P C by a singular shaft.

Description

T h e 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 L P T C a s e .

0 Inner and Outer Duct.

0 LPTShaf t .

0 Turbine Exhaust Case (TEC).

T h e LP turbine has a five stage rotor that supplies power to the LP compressor through the LPT shaft.

T h e LPT rotor is installed in the LPT case where it is in alignment with the LPT stators. T h e LPT case is made from high-heat resistant nickel alloy and is a one part welded assembly.

T h e LPT case has two borescope inspection ports found on the left and right hand sides. T h e ports are used to internally examine the adjacent engine sections:

Trailing Edge (TE), Stage 2, HPT Blades.

0 Leading Edge (LE), Stage 3, LPT Blades.

T h e remaining stages of the LP system do not have borescope access ports.

T h e five LPT disks are made from high heat resistant nickel alloy.

T h e LPT blades are also made from nickel alloy and are attached to the disks by firtree roots. T h e 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.


IAE V2500 Borescope Practices 0 IAE lnternalional Aero Engines AG 2000


N E 0

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 housing of 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 Individually replaceable 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 Integrated Drive Generator (IDG).



OIL FILTER

Tm

- STARTER LOCATION

/ HYDRAULIC PUMP

LOCATION


OIL TANK

FUEL PUMP DRIVE PAD

OIL SCAVENGE PUMP

PNUEMATIC

OIL PRESSURE PUMP

INTEGRATED DRIVE GENERATOR

LOCATION

HIGH SPEED GEARBOX AND ANGLE BOX

TlNG


BORESCOPE REQUIREMENTS

0 IAE International Aero Engines AG 2000 V2500 Borescope Practices Borescope Equipment

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 allowable damage.

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 dependant upon 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 periodicity checks.

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 INTENTIONALLY BLANK

Initial issue

0 IAE International Aero Engines AG 2000 V2500 Borescope Practices Borescope Equipment

Page 2-2

V2500 Borescope Practices 0 IAE International Aero Engines AG 2000

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 stand a rd eq u i pme n t . 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 Flexible fibrescope (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 150 watt 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 1 10-240 volt AC 50-400Hz power.

For portability some units operate from 12-15 volts DC s U P PlY 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

V2500 Borescope Practices Q IAE International Aero Engines AG 2000

Borescope Equipment

STEADY HANDLE \

STORAGE CASE

io

sa

LIGHT GUIDE CABLE

LIGHT SO

BOROSCOPES

STORAG-E CASE

BORESCOPE INSPECTION EQUIPMENT Initial issue Page 2-4

V2500 Borescope Practices

V2500 Borescope Equipment

Rigid Borescopes

0 IAE international Aero Engines AG 2000 Borescope Equipment

Fibrescopes

F o r the inspection of the V2500 engine rigid borescopes Fibrescopes are necessary if a full AMM inspection of the are the best recommended for use. engine is to be carried out. Fibrescopes have the

advantage of the reaching areas of the engine that rigids T h e following table shows the recommended scopes: cannot.

T h e 6mm and 8mm fibrescopes are recommended for Diameter Working length Direction of Field of view degs use. This recommendation is in conjunction with rigid mm 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.

T h e interchangeable tip has the same idents except the working length is replaced with:

AIOS

A10 field of view.

S direction of view.

s the cope

Page 2-5

Q IAE International Aero Engines AG 2000 V2500 Borescope Practices

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 PROBE TIPS

FLEXIBLE INSERTION TUBE

LIGHT GUIDE

RIGID BORESCOPES AND FIBRESCOPES



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 t issue 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.

0 perat i on

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.





Fi brescopes

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 borescope and 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.


0 IAE International Aero Engines AG 2000 V2500 Borescope Practices

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.

Inspect the 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 Subject the instrument to any unnecessary force.

0 Use in a flammable atmosphere.

0 Bend or hit the shaft.

0 Force the focusing barrel against the stops.

0 Force the rotational control against the stops.

0 Introduce the 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.



Care of Use of Borescope Equipment

Fi brescopes

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 protective caps 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 Twis t the bending section by hand.

0 Insert into live electrical equipment.

0 U s e in corrosive fluids.

0 Leave the finger on angle knobs when removing the scope.

0 U s e hard cloth or brush for cleaning.

0 Apply excessive force to the bending section.

0 Apply excessive force when inserting or removing the in s e rtion 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

as ses s 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) 1 loo retro b) 4 5 O fore-oblique c) 90° lateral

covers



MAINTENANCE PRACTICES


Maintenance Practices


T h e safety of personnel and the aircraft is of paramount importance. T h e maintenance practice section looks at the actions that can be carried out which safely allow a borescope of the engine to take place.

T h e 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 LP and HP shafts.

Igniter plug removal and installation.

Borescope access plugs removal/installation.

Aircraft Preparation

T h e preparation of the aircraft requires that a check in the flight deck be carried out.

T h e check is to ensure that the FADEC system power switch is in the off position. T h e 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 E N G panel 11 5VU.

Ensure that the engine has not been shut down within the last five minutes.


*I.. --. ..



ENGINE CONTROL PANEL-Vi 15

FLIGHT DECK Initial issue Page 3-2


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 H C U 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 of this when working on the engine immediately after shut down.

HCU Deactivation (AMM 78-30-00-040-01 2)

0 Carry out the flight deck checks a s per aircraft preparation.

0 Open the fan cowl doors (71-1 3-00-01 0-01 0).

0 Position the lock lever on the H C U 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 78-30-00-440-01 2) 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.



0 IAE International Aero Engines AG 2000 V2500 Borescope Practices Maintenance Practices

HYDRAUL

- HYDRAULIC CONTROL UNIT (HCU) DEACTIVATION AND U

REACTIVATION initial issue Page 3-4


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 le s s 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 -1 3-00-01 0-01 0

Carry out the flight deck checks a s 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 -I 3 - 0 0 4 0-01 0

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. 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 71-0259 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.



F



Thrust Reverser C Ducts Maintenance

Warning

T h e 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. 78-32-0041 0-01 0

Carry out the flight deck checks a s 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 -1 3-00-01 0-01 0).

Allow the thrust reverser units to close.

Note:

T h e forward most latch must be in the locked position before closing.

Deactivate the H C U (78-30-00-040-01 2).

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 1 through 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 H C U (78-30-00-01 0-01 0)

Close the fan cowl doors (71-1 3-00-410-01 0).

Return the aircraft back to its usual condition.


0 IAE international Aero Engines AG 2000 V2500 Borescope Practices Maintenance Practices

BIFURCATION

HYDRAULIC HAND

s 8 r In

t n THRUST REVERSER C DUCT OPENlNGlCLOSlNG Initial issue Page 3-8



HP System rotation point access

Removalllnstallation access cover

T h e access cover is located on the front face of the external gearbox between the air starter and dedicated alternator (A I /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.

T h e 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, a s required. Viewed from the front of the engine.

On completion of inspection, remove socket, lubricate with engine oil and fit new packing to cover.

Refit cover to gearbox and torque load to the specified values in the manual.

Automatic

Rotator K i t Part No. IAE 2F10057 or 2F10066.

T h i s unit provides a compressed air powered device to rotate the HP system through the access point on the gearbox. T h e unit has foot pedal control of both direction of rotation and speed of rotation; it also incorporates a protractor to give an indication of rotation in degrees.

Alternative Electronic Units

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 3-1 0

0 IAE lnternalional Aero Engines AG 2000 V2500 Borescope Practices Maintenance Practices

Igniter Plugs Maintenance

Removalllnstallation procedure

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 Maintenance Manual

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 -400-01 0

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 IGNITER LEAD

E 0 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

A I 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 0 IAE International Aero Engines AG 2000


2.5

LP COMPRESSOR BOOSTER BORESCOPE ACCESS Page 3-14 Initial issue


HPC Borescope Access

There are two SB standards of H P C designs for A I engines where the borescope access features differ. T h e following is an explanation of the pre and post modification stand a rds . A1 Engines pre SB 72-0033

T h e following information relates to engines that are prior to SN V0127.

T h e inspection requirements a s advised by the AMM for this standard of engine is:

0 Recommended that only stage 3 and stage 12 of the H P C blades be inspected while the engine is on wing.

0 Recommended not to use access port D a s possible damage can be caused to the inspection equipment.

0 A limited number of engines prior to SN V0063 will have H P C 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

T h e borescope access ports give access to these stages of the compressor:

Port A H P C stage 3 rear. It is located 34 deg below the engine horizontal on the right side.

Port B H P C stage 3 rear and stage 4 front. It is located 34 deg below the engine horizontal on the left and right side.

Port C H P C stage 5 rear and stage 6 front. It is located 67 deg below the engine horizontal on the right side.

Port D H P C stage 7 rear and stage 8 front. It is located 67 deg below the engine horizontal on the right side.

Port E H P C stage 8 rear and stage 9 front. It is located 74 deg below the engine horizontal on the right side.

Port F H P C stage 9 rear and stage 10 front. It is located 61 deg below the engine horizontal on the right side.

Port G H P C stage 11 rear and stage 12 front. It is located 55 deg below engine horizontal on the right side.

Note:

T h e access to the borescope plug C requires the removal of the control rod from the unison ring at the stator s i x VSVs. It is recommended that only the stage 3 and stage 12 HP compressor blades be examined with the engine on-wing.



PREMOD SB 72-01 00

PORTS E, F, G POST MOD SB 72-01 00 PORT D POST MOD SB 72-0100

SB 72-0100

PORT D PRE MOD PORT C


> & PORT b A

HPC BORESCOPE ACCESS A I PRE SB 7210027 Initial issue Page 3-16



A I Engines post 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 33.5 deg below the engine horizontal on the left side.

Port D HPC stage 7 rear and stage 8 front. It is located 67 deg below the engine horizontal on the right side.

Port E HPC stage 8 rear and stage 9 front. It is located 74 deg below the engine horizontal on the right side.

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 H P C stage 11 rear and stage 12 front. It is located 55 deg below engine horizontal on the right side.

Port H HPC stage 3 front (additional access). It is located 41 deg below the L.H split 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 s ix 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 recommended by 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 PORT A

PORT B

PORT C SB 72-0338

PORT D PORT C

SB 72-0265

3 N 0 In

[3 HPC BORESCOPE ACCESS A I POST SB 72-0027 Page 3-18 Initial issue


A5/D5 Engines HP Compressor Borescope Access

The borescope access ports give access to these stages of the compressor:

Port A HPC stage 3 front. It is located 40 deg below the engine horizontal on the left side.

Port B H P C stage 3 rear and stage 4 front. It is located 34 deg below the engine horizontal on the right side.

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 deg below the engine horizontal on the right side.

Port E HPC stage 8 rear and stage 9 front. It is located 74 deg below the engine horizontal on the right side.

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 11 rear and stage 12 front. It is located 55 deg below engine horizontal on the right side.

Note:

It is recommended that only the stage 3 and stage 12 HP compressor blades be examined with 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 compound enter the borescope access port hence into the engine. Initial issue


Page 3-1 9


-

PORT E F G


PORT C PORT C

SE 72-0317


PORT A

B PORT B

PORT C

HPC BORESCOPE ACCESS AWD5 Initial issue Page 3-20


HP Compressor Access 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.


Q IAE International Aero Engines AG 2000 V2500 Borescope Practices Maintenance Practices

CON1

CENTRALISING PAD

/ VSV OPERATING LEVER

S TAGE 5 UNISON RING

BORESCOPE ACCESS PORT C

’ HP COMPRESSOR BORESCOPE ACCESS PORT C Initial issue Page 3-22



Combustor, HP and LP Turbines Borescope Access

Borescope access for the combustor is found in eight positions, of which s i x 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 B1 to 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 B1 to 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:

T h e borescope access ports are located near the diffuser case rear flange. T h e ports must not be confused with the 5 larger locating pins that are equi spaced around the forward end of the case.

HP Turbine

T h e HP turbine has provision for inspection of the leading and trailing edges of the blades.

LP Turbine

T h e 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 a s 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.


Q IAE International Aero Engines AG 2000 Maintenance Practices V2500 Borescope Practices

<

I / 8

TRAILING EDGE HPT 1 LEADING EDGE HPT 2 COMBUSTION 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


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 description from 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 less than 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 continue forward 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.



Borescope Terms 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


0 IAE International Aero Engines AG 2000 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 acceptheject information 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 appropriate chapter.

The following inspection/check areas of the engine will be discussed in this section;

0 L P C booster.

0 HP compressor.

0 Combustion system.

0 HP turbine.

0 LP turbine.

Note:

Engine rejection due to damage found being out of limits must be reported to the local I A E representative.

Any damage found not covered in the AMM must be reported to the local I A E representative for further advice.

Inform your local I A E 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.

If the damage exceeds the ultimate limit then the engine is scheduled for removal as per AMM requirements.


V2500 Borescope Practices 0 IAE Internatlonal Aero Engines AG 2000


IAE V2500 PROPULSION UNIT Initial issue Page 4-4

Q IAE international Aero 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 I A E 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 a s follows:

Stage 1.5 52 blades 58 blades.

A I engines A5/D5 engines

Stage 2.0

Stage 2.3

68 blades 78 blades.

88 blades. ------------- Stage 2.5 70 blades 72 blades.

Inspection Standards

T h e L P C booster can be inspected while the engine is on wing.

T h e aircraft and engine must be in the prepared condition a s 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.


0 IAE International Aero Engines AG 2000 V2500 Borescope Practices Engine Inspection/Check

2.5

LP COMPRESSOR BOOSTER INSPECTION initial issue Page 4-6



LPC Booster Inspection

The images below show how access is gained to the stages of the L P C 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 Ex i t 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 1.5 blades 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

- \ \ GUIDE TUBES P

1.5 7ON

AlIA51D5 STAGE NSPECl

LPC BOOSTER INSPECTION Initial issue Page 4-8



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.

T h e three zones are identified as:

0 ZoneA.

0 ZoneB.

0 ZoneC.

T o help the engineer to as ses s damage according to which zone it is in the blade has been apportioned a percentage of the total aerofoil surface.

T h e 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 as ses s damage in the zones according to AMM recommendations the blade dimensions must be known.

T h e 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.

A I engines stage 2.0 cannot be accessed for borescope inspection.

T h e width is the true chord measurement.

T h e 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 4 STAGE 2.3 A5/D5

2

7

1 - ZONE C

ZONE B 50% OF

AEROFOIL ZONE A 4

STAGE 2.5 Al/A5/D5

LPC BLADE DIMENSIONS AND 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 s i ze is 6.0mm. 0 Rigid scope s i zes 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

A I engines A5/D5 engines

Stage 8

Stage 9

Stage I O

Stage 11






T h e 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 a s 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 1 shingling, (snubber override).

0 Examine for ceramic coating loss.

Any damage found can be assessed according to the AMM recommendations.





I \ /

I MID HEIGHT SUPPORT VORTEX REDUCER Rl

hl 0 v)

HP COMPRESSOR STAGES 3 TO 12 8 n




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 a s 4 separate zones.

T h e acceptlreject limits for damage found in these zones may vary between them.

T h e zones are identified as:

0 ZoneA.

0 ZoneB.

0 ZoneC.

H P C stage 3 has an extra zone known as Zone D.

T o help the engineer to as ses s damage according to which zone it is in the blade has been apportioned a dimensional s i ze and/or percentage of the total aerofoil surface.

HPC Stage 3

T h e H P C stage 3 blade is apportioned into four zones. They are a s 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).


V2500 Borescope Practices 0 IAE Internalional Aero 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 Initial issue Page 4-14


HP Compressor Inspection Blade Dimensions 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.

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 1 0,9 11,l

Note: These dimensions are for reference purposes only.

The chord dimension changes from the blade root to the blade tip.

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


. .


LEADING EDGE

f


HP COMPRESSOR BLADE DIMENSIONS Initial issue Page 4-1 6


HP Compressor Inspection

HPC Blade Tip Rubbing

T h e H P C 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.

T h e 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

H P C 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.

T h e damper wires of pre modification SB 72-0289 HPC's have exhibited damper wire detachment. T h i s can lead to secondary engine damage of the HPC.

SIL 106;

SB 72-0289 introduced a fix for the existing style of H P C disc that modifies the existing wires with rounded ends.

SB 72-0304 introduces a new type of wire. T h e 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.

U s e borescope access ports C, D and E for examination for missing wires.


acceptable as long a s 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 P b to the EEC.

T h i s 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

T h e 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 H P C stages.


. ., .-. . . . .. . . . . .


MATERIAL LOSS


STATOR BLADE

\ MID CHORD TIP RUB

\

CERAMIC COATED ABRAIDABLE STATOR PATH

2 st n HPC TIP RUB AND DAMPER WIRE DETACHMENT



Combustion Chamber and HPT Stage I NGV'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 req u i re men t 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:


Borescope access

Access for combustion chamber inspection is as follows:

The igniter plug port positions of IP1 and IP2.

Combustion chamber ports B1 through to B5 (Pre SB 72-

Combustion chamber ports B1 through 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

022 1 ).

0221).

commencing engine 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 1 to 5 starting at the FSN's.

The AMM has printed forms for the recording of damage found.




COMBUSTION CHAMBER BORESCOPE VIEWS Initial issue Page 4-20




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 Coating damage.

0 Connected loop crack.

Note:

Damage assessment can be greatly enhanced by using parts within the combustion system that are of a known dimension.

U s e parts that are of a known dimension that are nearest to the damage.

Special Inspection to SB 72-0349 T h i s 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.




COMBUSTOR OUTER LINER SHELL

I

OUTER LINER SEGMENT

' COMBUSTION CHAMBER AREAS OF INSPECTION Initial issue Page 4-22

Q IAE international Aero Engines AG 2000 V2500 Borescope Practices 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 inspection for 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 Page 4-23



TYPICAL VIEWS OF END CAP DAMAGE

LOSS RADIAL CRACKS \ CIRCUMFERENTIAL

CRACKS

CAP FUEL SPRAY

BULKHEAD SEGMENT FUEL SPRAY NOZZLE

FUEL SPRAY NOZZLE AND BULKHEAD SEGMENT Page 4-24 Initial issue



Combustor Inspection T h e inspection requirements for the combustor in the Note: AMM ask for the inspection of the 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 adhered to.

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 I i ners. 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.

Page 4-25



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 a require the following areas to be inspected: 0 Concave aerofoil surface. decrease borescope inspection 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

2.100 in (53,34 mm) (43,18 mm)

PLATFORM LEADING EDGE COOLING

AIR HOLES

HPT STAGE 1 VANE 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/n V0362.

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 FLEX1 SCOPE I

No 4 BEARING SUPPORT,

COOLING HOLE

No 4 BEARING

m w


loTH STAGE MAKE U

STAGE 10 MAKE UP AIR VALVE OFF TAKE TUBES

I- \ ENTRY POINT FOR

INSPECTION

METERING PLUGS LO CAT1 ON

INSPECTION OF THE HPT STAGE I METERING PLUGS Initial issue Page 4-30



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

T h e HPT is a 2 stage turbine system designed to drive the HPC.

Borescope access to the HPT is possible for both stages.

T h e 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.

T h e borescope access ports are located on both sides of the core engine. They are identified as follows:

HPT Stage 1 Leading Edge

0 IP1 and/or IP2.

0 B1 through to B5 (Pre SB 72-0221).

0 BI through to B6 (Post SB 72-0221).

HPT Stage 1 Trailing Edge, Stage 2 Leading Edge

0 T1/2L which is on the left hand side.

0 T1/2R which is on the right hand side.

HPT Stage 2 Trailing Edge

0 T2/3L that is the left hand side.

0 T 2 / 3 R that is the right hand side. Initial issue


Each stage of the HP turbine has a number of blades per disc. These are a s follows: 0 HPT stage I has 64 blades.

0 HPT stage 2 has 72 blades.


Inspect the HPT for 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.

Page 4-31



FUEL SPRAY NOZZLE

HPT 1 LIE-FLEX1 ACCESS THROUGH

IGNITER PORT

HPT 1 UE-RIGID L/E=LEADING EDGE ACCESS THROUGH T/E=TRAILING EDGE COMBUSTOR CASE

00 HPT 2 LIE 8 m

v)

HPT STAGES I AND 2 BORESCOPE VIEWS 2 W 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 between them.

The zones are identified as:

0 ZoneA.

0 ZoneB.

0 ZoneC.

0 ZoneD.

T o help the engineer to assess damage according to which zone it is in the blade has been apportioned a dimensional s ize and/or percentage of the total aerofoil 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 HP turbine stage 1 blade overall radial length is:

0 100% equals 1.80in (45,72mm).



0.187 in (4,75 mm)

t NOTE: 100% SPAN EQUALS

1.70 in (43,18 mm)

DEV250239

ZONE D

1

-

100%

STAGE 1 HPT BLADE CONVEX AIRFOIL VIEW STAGE 1 HPT BLADE CONCAVE AIRFOIL VIEW

HPT STAGE I BLADE ZONES Initial issue Page 4-34

0 IAE Internalionat Aero Englnes AG 2000 V2500 Borescope Practices

HP Turbine Stage 1 Blade Inspection

T h e 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 HP turbine blades for the following conditions:

Cracks, Nicks, Burns and Dents on the Airfoil Surface

Generally cracks, nicks, burns and dents are acceptable as long a s 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 a s long a s the AMM l imits 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 1 HPT blades some areas of the stage 1 HPT duct segments can be seen.

Engine InspectiorKheck



ROOT PLATFORM CRACK 0 STAGE 2 NGVs BLADE PLATFORM 8 0 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 may vary between them.

The areas in question are called zones. The zones are identified as:

0 ZoneA.

ZoneB.

0 ZoneC.

0 ZoneD.

T o help the engineer to assess damage according to which zone it is in the blade has been apportioned a dimensional s ize and/or percentage of the total aerofoil 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 Page 4-37

V2500 Borescope Practices 0 IAE lnternatlonal Aero Engines AG 2000


AXIAL DIRECTION

Initial issue

75%

50%

HP TURBINE STAGE 2 BLADE ZONES

Page 4-38



' HP Turbine Stage 2 Blade Inspection

T h e 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. T h i s will be the acceptable practice until the ultimate limit has been reached which will then require the engine to be removed.


Inspect the HP turbine 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 a s 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 a s long a s the AMM limits and procedures are adhered to.

Stage 1 HPT Duct Segment for Burn Holes

Generally burn holes are acceptable a s 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

HPT STAGE 2 NGVs

hl e In 8

Q E

\ BLADE ROOT PLATFORM

HPT STAGE 2 BLADE AND DUCT SEGMENT Initial issue Page 4-40



LP Turbine Stage 3 InspectionlCheck

AMM ref. 72-00-00-200-012 T h e LPT stage 3 rotor blade leading edge is accessible for inspection. There is a borescope port provided.

T h e LPT stage 7 trailing edge is accessible for inspection. This is viewed by looking down the exhaust nozzle.

F o r LPT stage 3 borescope access port T112L and T1/2R can be used.

T h e recommended inspection equipment to effect a borescope inspection of the LPT stage 3 is a s 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.

Inspection Standards T h e inspection standards for the LPT stage 3 are a s follows:

0 Cracks. 0 Nicks.

0 Tears.

0 Dents. Initial issue

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.

Page 4-41


LPT STAGE 3 STATOR VANE I //I

LPT

BORESCOPE PROBE

STAGE 3 ROTOR BLADE BORESCOPE

LPT STAGE 3 ROTOR BLADE

3

VIEW Initial issue Page 4-42

0 IAE lnternalional Aero Engines AG 2000 V2500 Borescope Practices

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 may vary between them.

The areas in question are called zones. The zones are identified as:

0 Z o n e A l .

0 ZoneA2.

ZoneA3.

T o help the engineer to assess damage according to which zone it is in the blade has been apportioned a dimensional s ize and/or percentage of the total aerofoil surface area.

Zone A I

Zone A I is O.l ins (2,5mm) from the blade shroud moving towards the root.

Zone A2

Zone A2 is the majority of the blade aerofoil surface.

T h i s 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 Borescope Practices @ IAE International Aero Engines AG 2000


\

A3

A1

A2

LEADING EDGE

0.120in (3,Omm) '

I

0.120in (3,Omm)

OUTER SHROUD

TRAILING EDGE ' 0.120in (3,Omm)

INNER SHROUD

BLADE ROOT

LP TURBINE STAGE 3 BLADE ZONES Initial issue Page 4-44


LP Turbine Stage 3 Blade Inspection

T h e AMM gives flexibility with the assessment of the LP turbine. As the damage deteriorates then the inspection interval will decrease and/or 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.


Inspect the LP turbine blades for the following conditions:

Cracks in areas A I A2 and A3

These are acceptable as long a s the limits in the AMM are adhered to.

Flame Plate Extrusion (area AI)

Generally this can be accepted from the Z notch as per AMM recommendations.

Dents and Nicks in areas AI , A2 and A3

These are acceptable as long as the limits in the AMM are ad hered to.

Tears and Bends in area A2

These are acceptable as long a s the limits in the AMM are adhered to.

Build Up of Deposits

T h i s can generally be accepted as per AMM recommendations. Initial issue

0 IAE international Aero Engines AG 2000 Engine InspectionlCheck

Blistering and Loss of Coating (all areas)

These are acceptable a s long a s the limits in the AMM are adhered to.

Sulphidation (all areas)

Sulphidation appears as a greenish to pale blue colour.

These are acceptable a s long a s the limits in the AMM are adhered to.

Burning and Oxidation on Outer shrouds

Generally accept burning and oxidation on side faces that do not touch as per AMM recommendation.

Examine Blanking Plug and Sleeve Flange Bushing

Reject if cracked or bent as per AMM recommendations.

Examine the plug for fretting a s per AMM recommendations.

Generally check for damage. Any damage that prevents the proper use of the plug reject a s per AMM recommendations.

Note:

U s e the spacer plate washers as removed before. This makes sure that the blanking plug has the same fit a s before a s long as the original fit is in limits. This is according to AMM recommendations.

Page 4-45


0

FLAME PLATE (OUTER SHROUD)

LPT STAGE 3 ROTOR BLADE

i LEADING EDGE

BLANKING PLUG


TRAILING EDGE

INNER SHROUD

LPT STAGE 3 ROTOR BLADE AND BLANKING PLUG Initial issue Page 4-46

CLASSIFICATION OF BLADE AND NGV DAMAGE


Classification of Blade and NGV Damage

Initial Issue

DAMAGE THAT IS CAUSED TO A PART THAT IS CONSTANTLY HIT

BATTERED Page 5-1

0 IAE International Aero Engines AG 2000 IAE V2500 Borescope Practices Classification of Blade and NGV Damage

Initial Issue

OVERHEATED Page 5-2

IAE V2500 Borescope Practices 0 IAE international Aero Engines AG 2000


A SMALL CUT ON THE SURFACE OR EDGE OF A PART. CAUSED WHEN THE PART IS HIT WITH AN OBJECT

NICK Initial Issue Page 5-3



Initial Issue

A LARGE ROUGH CUT OF LARGE DEPTH WITH SOME REMOVAL OF MATERIAL CAUSED BECAUSE

A SHARP OBJECT HAS HIT THE PART

GOUGE Page 5-4

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IAE V2500 Borescope Practices 0 IAE international Aero Engines AG 2000

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DEPOSITS

' % - [ ... :e==-- PARTICLES OF MATERIAL COLLECTED ON A PART,

FROM A DIFFERENT PART OR MATERIAL

Initial Issue Page 5-5

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0 m m 0 d ? r n w

U DAMAGE TO A SURFACE OF A PART WHEN IT IS HIT WITH AN OBJECT. THE MATERIAL

IS DISTORTED BUT NOT REMOVED

DENT Initial Issue Page 5-6



CURL Initial Issue

A ROUNDED FOLD IN THE MATERIAL SUCH AS A BLADE TIP THAT HAS RUBBED AGAINST THE ENGINE CASING

Page 5-7



A LINEAR OPENING THAT CAN EASILY BE SEEN AND WHICH CAN CAUSE THE MATERIAL TO BREAK

CRACK Initial Issue Page 5-8

0 IAE International Aero Engines AG 2000 IAE V2500 Borescope Practices Classification of Blade and NGV Damage

BURNED

COMPLETE STRUCTURAL FAILURE OF THE MATERIAL BECAUSE OF VERY HOT TEMPERATURES

Initial Issue Page 5-9



THE SEPARATION OF A PART BY FORCE IN TWO OR MORE PIECES

Initial Issue

BROKEN Page 5-10


BENT


ANGULAR CHANGE FROM THE INITIAL SHAPE OR CONTOUR. USUALLY THE

CAUSE IS A LATERAL FORCE

Initial Issue Page 5-1 1

TYPICAL EXAMPLES


1

cb IAE international Aero Engines AG 2000 Typical Examples

FUEL SPRAY NOZZLE END CAP DAMAGE Initial issue Page 6-1

Initial issue

cb IAE International Aero Englnes AG 2000 Typical Examples

BULKHEAD SEGMENT AND DEFLECTOR DAMAGE Page 6-2

&D IAE International Aero Engines AG 2000 V2500 Borescope Practices Typical Examples

FUEL SPRAY NOZZLE END CAP DAMAGE Initial issue Page 6 3

V2500 Borescope Practices r--- aB IAE international Aero Engines AG 2000

Typical Examples I n . - . . ”.. . .~

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COMBUSTOR LINER SEGMENT DAMAGE Initial issue Page 6-4

r V2500 Borescope Practices

Initial issue

(B IAE International Aero Engine8 AG 2OOO Typical Examples

STAGE 1 HPNGV DAMAGE Page 6-5

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V2500 Borescope Practices Q IAE Internatlonal Aero Englner AG 2000

Typical Examples

STAGE I HP TURBINE DAMAGE Initial issue Page 6-6


initial issue

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Q IAE International Aero Engines AG 2000 Typical Examples

STAGE I HP TURBINE BLADE DAMAGE Page 6-7

(D IAE IntemaUonal Aero Engines AG 2OOO

1 V2500 Borescope Practices

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- 1

Typical Examples

COMBUSTOR TILE SEGMENT DAMAGE Initial issue Page 6-8

V2500 Borescope Practices Q IAE Internatlonal Aero Englnes AG 2000

Typical Examples

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HPC PLATFORM DAMAGE Initial issue Page 0-9


Initial issue

Q IAE lnternatlonal Aero Englnes AG 2000

I

HP COMPRESSOR STAGE 7 DAMAGE

Typical Examples

Page 6-1 0

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Initial issue

0 IAE International Aero Engines AG 2000 Typical Examples

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VSV - OIL STAINING FROM NO 3 BEARING Page 6-1 1


Typical Examples

Initial issue

LINER SEGMENT VIEWED FROM UNDERSIDE Page 6- 12

0 IAE International Aero Engines AG 2000 Typical Examples

F

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HPT STAGE 1 BLADE TIP WITH CRACK Initial issue Page 6-1 3

Initial issue


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HPC ROTOR DRUM LINER DAMAGE

Typical Examples

Page 6-14

Q IAE International Aero Engines AG 2000 Typical Examples V2500 Borescope Practices

OUTER COMBUSTION 4 CHAMBER LINER DAMAGE Page.6-15 Initial issue

Technology

V2500 bsi issue 01