CF34-10E LM June 09 print.pdf

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

CF34-10E TRAINING MANUAL

Document: CF34-10E Revised: June, 2009

Published By: Customer Technical Education Center 123 Merchant Street Mail Drop Y2 Springdale, Ohio 45246

EFFECTIVITY ALL GE PROPRIETARY INFORMATION

INTRO INTRO

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CF34-10E LINE MAINTENANCE MAINTENANCE This publication publication is for TRAINING PURPOSES PURPOSES ONLY. ONLY. This information is accurate at the time of compilation; however, no update service will be furnished to maintain accuracy. For authorized maintenance practices and specifications, consult the pertinent Maintenance Manual. This product is considered GE Aircraft Engines technical data information and therefore is exported under U.S. Government Export License Regulations. It is issued to the user under specific conditions that the data, or it’s product may not be resold, diverted, transferred, transshipped, reexported, or used in any other country without prior written approval of the U.S. Government. The information contained in this document is disclosed in confidence. It is the property of GE Aircraft Engines and shall not be used (except for evaluation), disclosed to others, or reproduced without the expressed written consent of GEAE. If consent is given for reproduction in whole or in part, this notice shall appear on any reproduction, in whole or in part. The foregoing is subjected to any rights the U.S. Government may have acquired as such information.

Copyright 2005 GE Transportation Published by: GE Aircraft Engines Customer Training Services Customer Technical Education Center 123 Merchant Street Cincinnati, Ohio 45246 EFFECTIVITY ALL GE PROPRIETARY INFORMATION

INTRO INTRO

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TABLE OF CONTENTS Section INTRO 1 2 3 4 5 6 7 8 9 10 11

Topic Introduction INTRODUCTION POWERPLANT FAULT DETECTION STRATEGY ENGINE GENERAL FUEL AND CONTROL IGNITION SYSTEM ENGINE AIR SYSTEM ENGINE INDICATING ENGINE OIL SYSTEM ENGINE STARTING ENGINE EXHAUST


Revision Aug 05 Sept 05 June 09 Feb 09 June 09 June 09 Feb 09 Feb 09 June 08 June 09 June 08 Aug 06

INTRO INTRO

Page Numbers 1-5 1-13 1-19 1-23 1-23 1-59 1-17 1-15 1-27 1-31 1-19 1-47

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ACRONYMS AND ABBREVIATIONS A/C – AIRCRAF AIRCRAFT T ADC – AIR DATA DATA COMPUTER COMPUTER ADS – AIR DATA DATA SYSTEM SYSTEM AGB – ACCESSARY ACCESSARY DRIVE DRIVE GEARBOX AI – ANTI-IC ANTI-ICING ING A/I - APPROACH APPROACH IDLE IDLE AIP – AUTONOMOUS AUTONOMOUS INPUT PROCESSOR PROCESSOR ALF – AFT LOOKING LOOKING FORWARD FORWARD AMM – AIRCRAFT AIRCRAFT MAINTENANCE MAINTENANCE MANUAL APPID APPID – APPLICATION APPLICATION IDENTIFICATI IDENTIFICATION ON APR – AUTOMATIC AUTOMATIC POWER RESERVE RESERVE APU – AUXILIARY AUXILIARY POWER POWER UNIT ARINC – AERONAUTICAL AERONAUTICAL RADIO INC. INC. AS – APPLICATIO APPLICATION N SOFTWARE ASCB – AVIONICS AVIONICS STANDARD STANDARD COMMUNICATIONS BUS AT – AUTO THROTTLE THROTTLE ATA – AIR TRANSPORT TRANSPORT ASSOCIATION ASSOCIATION ATSV – AIR TURBINE TURBINE STARTER VALVE VALVE ATTCS – AUTOMATIC AUTOMATIC TAKEOFF TAKEOFF THRUST CONTROL SYSTEM

CL – COWL LOCK LOCK CMC – CENTRAL MAINTENAN MAINTENANCE CE COMPUTER CPU – CENTRAL PROCESSI PROCESSING NG UNIT CW – CLOCKWI CLOCKWISE SE DCU – DIRECTIONAL DIRECTIONAL CONTROL CONTROL UNIT D/I – DECENT IDLE DIS – DISCRETE DISCRETE DATA DATA EBU – ENGINE BUILD UNIT UNIT ECP – ENGINE CONFIGURATI CONFIGURATION ON PLUG ECS – ENVIROMENTAL ENVIROMENTAL CONTROL SYSTEM SYSTEM EDP – ENGINE DRIVEN DRIVEN PUMP PUMP EHSV – ELECTRO-HYDRAUL ELECTRO-HYDRAULIC IC SERVO VALVE EICAS EICAS – ENGINE INDICATION INDICATION AND AND CREW ALERT SYSTEM EICC – ESSENTIAL ESSENTIAL INTERGRATED INTERGRATED CONTROL CENTER EVM – ENGINE VIBRATION VIBRATION MONITORING MONITORING SYSTEM

FADEC – FULL AUTHORITY AUTHORITY DIGITAL ENGINE ENGINE CONTROL BIT – BUILT-IN-TE BUILT-IN-TEST ST FDR – FLIGHT DATA RECORDER RECORDER BSI – BORESCOPE BORESCOPE INSPECTION INSPECTION F/I – FLIGHT IDLE FIM – FAULT ISOLATION ISOLATION MANUAL MANUAL CAS – CREW ALERT ALERT SYSTEM SYSTEM FMU – FUEL METERING METERING UNIT UNIT CCD – COMPUTER CURSER CURSER DEVICE DEVICE FOD – FOREIGN FOREIGN OBJECT OBJECT DAMAGE CCDL – CROSS CHANNEL CHANNEL DATA LINK LINK FRM – FAULT REPORT REPORT MANUAL MANUAL CCW – COUNTER CLOCKWISE FWSOV – FIREWALL SHUTOFF VALVE VALVE CDP – COMPRESSOR COMPRESSOR DISCHARGE DISCHARGE PRESSURE FMV – FUEL METERING METERING VALVE VALVE CIT – COMPRESSOR COMPRESSOR INLET INLET TEMPERATURE TEMPERATURE


G/I – GROUND IDLE GMO – GROUND MAINTENAN MAINTENANCE CE OVERRIDE OVERRIDE GPH – GALLONS GALLONS PER HOUR GPU – GROUND POWER UNIT UNIT GSBIT – GROUND START BIT BIT HP – HIGH PRESSURE PRESSURE HPC – HIGH PRESSURE PRESSURE COMPRESSOR COMPRESSOR HPT – HIGH PRESSURE PRESSURE TURBINE TURBINE HPTACC – HIGH PRESSURE PRESSURE TURBINE TURBINE ACTIVE CLEARANCE CONTROL HX – HEAT HEAT EXCHANGER EXCHANGER HYD – HYDRAUL HYDRAULIC IC HZ – HERTZ (CYCLES (CYCLES PER SECOND) SECOND) ICC – INTERGRATED INTERGRATED CONTROL CENTER CENTER ICU – ISOLATION ISOLATION CONTROL UNIT UNIT IDG – INTERGRATED INTERGRATED DRIVE GENERATOR GENERATOR IGB – INLET GEARBOX GEARBOX IGN - IGNITI IGNITION ON IGV – INLET GUIDE GUIDE VANE VANE ITT – INTER TURBINE TURBINE TEMPERATURE TEMPERATURE L/I – LANDING LANDING IDLE IDLE LP – LOW PRESSU PRESSURE RE LPT – LOW PRESSURE PRESSURE TURBINE TURBINE LPTACC – LOW PRESSURE PRESSURE TURBINE TURBINE ACTIVE CLEARANCE CONTROL LRU – LINE REPLACEAB REPLACEABLE LE UNIT LVDT – LINEAR LINEAR VARIABLE DIFFERENTIAL DIFFERENTIAL TRANSFORMER

INTRO INTRO

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ACRONYMS AND ABBREVIATIONS MAU – MODULAR MODULAR AVIONICS AVIONICS UNIT MCD – MAGNETIC MAGNETIC CHIP DETECTOR DETECTOR MEL – MINIMUM MINIMUM EQUIPEMENT EQUIPEMENT LIST MFD – MULTI FUNCTIONAL FUNCTIONAL DISPLAY DISPLAY N1 – FAN SPEE SPEED D N2 – CORE CORE SPEED SPEED NAC - NACELL NACELLE E NAI – NACELLE NACELLE ANTI-ICE ANTI-ICE NDOT – CORE SPEED SPEED ACCELERATION ACCELERATION RATE NPBIT NPBIT – NORMAL PERIODI PERIODIC C BIT NVM – NON VOLATILE VOLATILE MEMORY MEMORY OAT – OUTSIDE AIR AIR TEMPERATURE TEMPERATURE OD- OUTSIDE DIAMETER DIAMETER OEI – ONE ENGINE ENGINE INOPERATIVE INOPERATIVE OGV – OUTLET GUIDE GUIDE VANE VANE OVRD OVRD – OVERRI OVERRIDE DE PS3 – COMPRESSOR COMPRESSOR DISGARGE DISGARGE PRESSURE P0 – AMBIENT AMBIENT PRESSURE PRESSURE PAL – PROGRAMMAB PROGRAMMABLE LE ARRAY LOGIC LOGIC PDO – POWER DOOR OPENER OPENER PMA – PERMANENT PERMANENT MAGNET ALTERNATOR ALTERNATOR PMAT – PORTABLE PORTABLE MAINTENANCE MAINTENANCE ACCESS TERMINAL PPH – POUNDS PER HOUR HOUR POBIT – POWER ON BIT PSROV – PRESSURE PRESSURE REGULATED SOLENOID SOLENOID OPERATED VALVE PSI – POUNDS PER PER SQUARE INCH INCH PSIA – POUNDS PER SQUARE SQUARE INCH ABSOLUTE ABSOLUTE

TCASE – CASING TEMPERA TEMPERATURE TURE TCQ – THROTTLE CONTROL CONTROL QUADRANT QUADRANT TDS – TAKEOFF DATA SET SET TGB – TRANSFERE TRANSFERE GEARBOX GEARBOX TLA – THROTTLE LEVER ANGLE TLD – TIME LIMITED LIMITED DISPATCH DISPATCH T/O – TAKE OFF TOGA – TAKE OFF OFF GO AROUND AROUND TQA – THROTTLE QUADRANT QUADRANT ASSEMBLY ASSEMBLY TRAS – THRUST REVERSER REVERSER ACTUATION ACTUATION SYSTEM TRF – TURBINE REAR FRAME FRAME TRS – THRUST RATING RATING SELECTOR SELECTOR

PSID – POUNDS PER PER SQUARE INCH INCH DIFFERENTIAL PSIG – POUNDS PER PER SQUARE SQUARE INCH GAGE QEC – QUICK ENGINE ENGINE CHANGE CHANGE QTY – QUANTI QUANTITY TY RDS – RADIAL DRIVE SHAFT SHAFT R/I – REVERSE REVERSE IDLE IDLE RTD – RESISTIVE RESISTIVE THERMAL THERMAL DEVICE DEVICE RPM – REVOLUTIONS REVOLUTIONS PER MINUTE MINUTE RVDT – ROTARY VARIABLE VARIABLE DIFFERENTIA DIFFERENTIAL L TRANSFORMER SDI – SOURCE DESTINATION DESTINATION IDENTIFIE IDENTIFIER R SFA – SYNCHRONIZED SYNCHRONIZED FEEDBACK FEEDBACK ACTUATOR SFC – SPECIFIC SPECIFIC FUEL CONSUMPTION CONSUMPTION SLA - SYNCHRONIZED SYNCHRONIZED LOCKING LOCKING ACTUATOR SOV – SHUT OFF OFF VALVE VALVE SPDA – SECONDARY SECONDARY POWER DISTRIBUTION DISTRIBUTION ASSEMBLY

VAC – VOLTS ALTERNATIN ALTERNATING G CURRENT VDC – VOLTS DIRECT DIRECT CURRENT CURRENT VBV – VARIABLE VARIABLE BLEED BLEED VALVE VIB - VIBRATI VIBRATION ON VG – VARIABLE VARIABLE GEOMETRY GEOMETRY VSV – VARIABLE VARIABLE STATOR VANE VANE

T12 – FAN TOTAL INLET INLET TEMPERATURE TEMPERATURE T25 – COMPRESSURE COMPRESSURE INLET INLET TEMPERATURE TEMPERATURE T3 – COMPRESSOR COMPRESSOR DISCHARGE DISCHARGE TEMPERATURE TEMPERATURE T495 – INTER TURBINE TURBINE TEMPERATURE TEMPERATURE TAI – THERMAL THERMAL ANTI-ICE ANTI-ICE TAT – TOTAL AIR AIR TEMPERATURE TEMPERATURE TAMB – STATIC AIR AIR TEMPERATURE TEMPERATURE (CALCULATED) TBV – TRANSIENT TRANSIENT BLEED BLEED VALVE TC - THERMOC THERMOCOUPL OUPLE E


WAI – WING ANTIANTI-ICE ICE Wf – FUEL FUEL FLOW FLOW WFX – FMU METERING METERING VALVE VALVE POSITION POSITION FEEDBACK WOW – WEIGHT WEIGHT ON WHEELS WHEELS WS – WHEEL WHEEL SPIN SPIN Z1BRG – NO. 1 BEARING BEARING VIBRATION VIBRATION SENSOR ZFFCC – FAN FRAM COMPRESSOR COMPRESSOR CASE VIBRATION SENSOR

INTRO INTRO

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INTRODUCTION


00-00-00 INTRODUCTION

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INTRODUCTION Objectives Given an objective exercise, the student will identify: - The CF34-10E engine (1.A.a) - The maintenance documents necessary to do line maintenance (1.A.a) - The maintenance documents ATA numbering (1.A.a) - Alternate troubleshooting resources (1.A.a)

Given an objective exercise, the student will select the purpose of: - The CF34-10E engine (2.B.b)

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ENGINE INTRODUCTION Identification (1.A.a) The CF34-10E engine combines the most advanced technology available today with nearly 500 million flight hours of experience gained on the entire GE product line, from the GE90, the world's largest, most powerful engine, to the CF6 and CFM56, the bestselling, most reliable engines for 100 plus passenger aircraft.

Purpose (2.B.b) The CF34-10E engine has been selected to supply thrust for the EMBRAER 190/195 regional jet. The CF34-10E engine also provide power to operate these systems: -Electric -Hydraulic -Pneumatic

Flat rated at takeoff up to 86*F (30*C) with a thrust range of 16,960 to 18,820 pounds for the ERJ 190 application and up to 20,360 for the ERJ 195 application, the CF34-10E baseline concept includes the proven, simple and rugged architecture of the CF34/CFM family.

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Key CF34-10E engine design features include: a widechord fan for higher thrust and high tolerance to foreign object damage; 3-D aerodynamic design airfoils in the high-pressure compressor, providing highly efficient, stall-free operation, as well as better fuel burn and higher exhaust gas temperature margins ; a highly durable single annular, low-emissions combustor that meets or surpasses the most stringent emissions standards ; and a single-stage high-pressure turbine for lower operating cost. Refer to figure, CF34-10E ENGINE.



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MAINTENANCE DOCUMENTS

Refer to figure, MAINTENANCE DOCUMENTS.

Identification (1.A.a) The maintenance documents for the aircraft supply help for all maintenance activities. Many different documents work together to help the operator to do maintenance on the aircraft. The maintenance documents will help the operator do unscheduled line maintenance and scheduled line maintenance. Each maintenance document has an introduction that shows how to use that document.

Scheduled Maintenance These are examples of scheduled line maintenance work: - Aircraft turn around - Aircraft daily checks - Planned checks (A, B, C and D checks)

Use these documents to do unscheduled line maintenance: - Structural Repair Manual (SRM) - Fault Reporting Manual (FRM) - Fault Isolation Manual (FIM) - Master Minimum Equipment List (MMEL) - Aircraft Maintenance Manual (AMM) Use these documents to do scheduled line maintenance: - Maintenance Planning Document (MPD) - Maintenance Task Card Manual - Aircraft Maintenance Manual (AMM) 6 0 0 1 5 0 0 0 0 0 0 0 D S

CF34-10E

Use these documents to get support data to do scheduled and unscheduled maintenance: - System Schematics Manual (SSM) - Wiring Diagram Manual (WDM) - Aircraft Illustrated Parts Manual (AIPC)


Unscheduled Maintenance These are examples of unscheduled line maintenance work: - Flight faults - Ground faults - Service problems - Structural damage Structural Repair Manual The SRM supplies descriptive information and specific instructions for field repair of aircraft structure. The SRM has data related to the topics that follow: - Evaluation of permitted damage - Typical repairs - Material identification - Material substitution - Fastener installation - Alignment check - Planning Fault Reporting Manual The FRM supplies the fault codes of aircraft faults to the


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flight crew. These faults can be flight compartment effects or other faults. The FRM has standard logbook write-ups for each fault code. The fault codes refer the operator to the FIM.

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The SDS has a description of the interfaces, function, and operation of the systems and subsystems of the aircraft.

Fault Isolation Manual Maintenance crews use these resources to identify the maintenance procedures (FIM task numbers) to correct faults: - Crew observations - CMC generated maintenance messages - Fault codes and descriptions from the FIM - Front face bite equipment

The practices and procedures sections have data related to the functions that follow: - Maintenance practices - Servicing - Removal and installation of components - Adjustment and test - Inspection and check - Cleaning and painting - Repair

The FIM also has related references to the AMM and gives procedures necessary to verify when faults have been corrected.

Maintenance Planning Document The maintenance planning document specifies the tasks for each type of scheduled maintenance check.

Master Minimum Equipment List The MMEL supplies the recommended minimum equipment necessary for dispatch. MMEL procedures are located in Part II of the AMM. The MMEL also supplies the procedures for dispatch with a fault, if permitted. The maintenance crew uses the MMEL to decide when to fix the fault.

Maintenance Task Card Manual The maintenance task card manual has the task cards the operator uses during the maintenance checks.

Aircraft Maintenance Manual The maintenance crew uses the AMM to do the applicable maintenance procedure. The AMM has two parts: the system description section (SDS) and the practices and procedures section.


System Schematics Manual The system schematics manual helps the user to understand the operation of the system and to do a fault isolation procedure. The SSM supplies the interconnection of all LRUs of a system or subsystem. Wiring Diagram Manual The wiring diagram manual supplies details of the point-to-point wiring on the aircraft.


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Standard Practices Manual The standard practices manual has standard instructions applicable to necessary maintenance and repair of the aircraft. Aircraft Illustrated Parts Catalog The aircraft illustrated parts catalog has figures and related parts lists to help the operator identify the necessary replacement parts and their data. The applicable data includes the items that follow: - Figure that shows the location of the part - Replacement part number - Replacement part quantity - Supplier data - Specification numbers - Recommended spares - Service bulletin history

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MAINTENANCE DOCUMENTS ATA NUMBERING Identification (1.A.a) The aircraft manual elements are labeled in accordance with Air Transport Association (ATA) of America Specification 100. The ATA number system lets the user uniquely identify a component. ATA Number System The Aircraft Maintenance Manual (AMM) is divided into chapters, which supplies a functional breakdown of the complete aircraft. The chapters of the AMM are further divided by a section number. The sections are divided by a subject number. The three-element chapter sectionsubject-number (XXYYZZ) is an indicator that lets the user identify a single functional item. Each of the three elements of the indicator have two digits. The chapter number (first element) and the first number of the section number (second element) are assigned by ATA Specification 100. The subject number (third element) identifies individual subject designations (engine modules and piece parts) of the section of each chapter. 0 1 0 1 5 0 0 0 0 0 0 0 D S

ATA chapter numbers recorded in the CF34-10E engine program: - 70, Standard Practices - 71, Powerplant - 72, Engine General - 73, Engine Fuel and Control System - 74, Ignition System - 75, Air System - 77, Engine Indicating System - 78, Exhaust System - 79, Lubrication System - 80, Starting System Refer to figure, MAINTENANCE DOCUMENT ATA NUMBERING.

Material which is applicable to a system as a whole uses zeros in the second and third element of the number that is the chapter number followed by "-0000". An example is 72-00-00 (Engine) is given for general description data which supplies an outline breakdown of the section in the chapter. The list that follows is a description of the



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ALTERNATE TROUBLESHOOTING RESOURCES Identification (1.A.a) The data that follows is given for additional troubleshooting resources: Refer to figure, ALTERNATE TROUBLESHOOTING RESOURCES. Purpose (2.B.b) Identifies resources for the customer to get additional data on maintenance practices or procedures.

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[email protected] 3 1 0 1 1 0 0 0 0 0 0 0 D S



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POWERPLANT


71-00-00 POWERPLANT

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POWERPLANT Objectives Given an objective exercise, the student will identify: - Selected engine ratings and aircraft applications of the CF34-10E engine (1.A.a) - The location and information on the engine data plate (1.A.a) - Selected specifications of the CF34-10E engine (1.A.a) - Selected ground transportation requirements (1.A.a) - Selected engine safety hazards (1.A.a) - Selected components of the engine cowling (1.A.a) - Air management system (1.A.a) - Thermal anti-ice system (1.A.a) - Engine fire zones (1.A.a) - Fire detection system (1.A.a) - The location of the engine drains (1.A.a) - The engine mounts (1.A.a)

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Given an objective exercise, the student will select the purpose of: - The engine data plate (2.B.b) - Selected specifications of the CF34-10E engine (2.B.b) - Selected ground transportation requirements (2.B.b) - Selected components of the engine cowling (2.B.b) - Air management system (2.B.b) - Thermal anti-ice system (2.B.b) - Engine fire zones (2.B.b) - Fire detection system (2.B.b) - The engine drains (2.B.b) - The engine mounts (2.B.b)


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RATINGS AND APPLICATIONS Identification (1.A.a) The CF34-10E engine will be used on EMBRAER 190/195. There are different ratings for different variants of the CF34-10E engine: • • • • • •

CF34-10E2A1 CF34-10E5 CF34-10E5A1 CF34-10E6 CF34-10E6A1 CF34-10E7

Refer to figure, ENGINE RATING.

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CF34-10E2A1

CF34-10E5

CF34-10E5A1

CF34-10E6

CF34-10E6A1

CF34-10E7

86°F

86°F

86°F

95°F

95°F

86°F

(30°C )

(30°C )

(30°C )

(35°C )

(35°C )

(30°C )

Normal Takeoff Thrust

16,960

17,930

18,820

17,390

18,820

18,820

Max Takeoff (APR),

16,960

18,820

18,820

18,820

18,820

20,360

Sea Level Flat Rating

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ENGINE RATING


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DATA PLATE Identification (1.A.a) The engine data plate is a small, square metal plate on the fan case at the 9 o’clock position, aft of the Accessory Gearbox. Refer to figure, DATA PLATE. Purpose (2.B.b) The engine data plate records the following data: - Engine type and model - Production codes - Serial number - Manufacturer - Thrust rating - General engine data - Regulating agency data - Engine performance data - N1 trim information - Service bulletin compliance

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Engine Left Side

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DATA PLATE


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SPECIFICATIONS Identification (1.A.a) The CF34-10E engine specification are as follows: Length, Fwd spinner-to-aft CB (in./meters) Maximum Diameter (in./meters) Fan Diameter (in./meters) Weight, Dry with Mounts (lbs./kg) Core Speed Redline (RPM) Fan Speed Redline (RPM) Fan Bypass Ratio (at SLS Max) Overall Pressure Ratio (at SLS Max)

146.1/3.711 80.9/2.055 53/1.346 3675/1667 18018 6325 5.4:1 28.5:1

Purpose (2.B.b) The engine specification table shows the engine length, width, height, and weight of the CF34-10E engines. Refer to figure, CF34-10E ENGINE SPECIFICATIONS.

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GROUND TRANSPORTATION REQUIREMENT Identification (1.A.a) The engine can be transported by truck that has a pneumatic suspension or by air. These guidelines should be followed when the engine is moved: - Install the engine on a stand that agrees with GE specifications. - Use correct tie-down points to install the engine stand on the truck. - When a single engine is moved, install it over the trailer axle. The trailer must have pneumatic suspension. - If multiple engines are moved, the tractor and trailer must both have pneumatic suspensions. - Use the correct tie-down points to install the engine stand base on the truck. Always use the shipping stand shock mounts. Refer to figure, GROUND TRANSPORTATION REQUIREMENT. 0 1 0 1 5 0 0 0 0 0 1 7 D S

Purpose (2.B.b) Proper transportation of the engine is necessary to prevent shocks and vibration that can damage the engine bearings.


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ENGINE SAFETY HAZARDS Identification (1.A.a) The four engine safety hazard areas that should be avoided during engine operation are: - Inlet suction - Engine heat - Exhaust velocity - Engine noise Inlet Suction Inlet suction can pull people and large objects into the engine. At idle power, the inlet hazard area is a 13 ft (4.0m) radius around the inlet. WARNING: If the wind is more than 25 knots, the inlet hazard area increases by 20%.

engine in operation. Refer to figure, ENGINE SAFETY HAZARDS. Training Information Points (3.E.e) When the engine is in operation, the anticollision lights are usually on. Personnel must not go near an engine in operation except when: - The engine is at ground idle - The person operating the engine from the flight deck can communicate with the ground personnel The entry/exit corridors are between the inlet hazard areas and the exhaust hazard areas.

Engine Exhaust The engine exhaust is very hot for long distances behind the engine. The heat hazard area can cause damage or injury.

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Exhaust Velocity The exhaust velocity is very high for long distances behind the engine. The exhaust velocity hazard area can cause damage or injury. Engine Noise Engine noise can cause temporary and permanent loss of hearing. Ear protection must be worn when near an


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13 FEET AT IDLE

55’ IDLE

474’ TO

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65 MPH OR GREATER EXHAUST VELOSITY


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ENGINE COWLING Identification (1.A.a) The engine cowling has the following cowls: - Inlet cowl - Fan cowl - Thrust reverser cowl - Core cowl Refer to figure, ENGINE COWLING. Inlet Cowl The inlet cowl is a fixed interchangeable aerodynamic fairing, which supplies the inlet airflow to the fan and core sections of the engine. It is mounted on the forward face of the engine fan case and isolated from the engine core cowling. The assembly is composed of an inlet lip, forward bulkhead, an outer barrel, an acoustic inner barrel, an aft bulkhead, and an aft flange. The inlet cowl assembly includes ant-ice ducts, T12 sensor, and provisions for the FADEC and fan compartment ventilation.

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Fan Cowl The fan cowl door assemblies are engine-to-engine interchangeable units enclosing the engine fan case between the inlet cowl and the thrust reverser cowl. Each assembly is supported by three hinges at the pylon and latched to the other fan cowl along the bottom split line with three tension hook latches. A hold open rod supports each door in the open position. When opened,


these cowls provide access to all hardware mounted on the outer fan case. Thrust Reverser Cowl The thrust reverser cowl is a bifurcated assembly of two halves. The fixed structure wall and the translating cowl form the fan exhaust duct and nozzle. The inner wall of the fixed structure, along with the core cowl, encloses the engine between the fan frame and the nozzle, and provides a fireproof boundary around the fire zone constituted by the engine. The thrust reverser contains a hydraulically powered thrust reverser actuation system (TRAS) to reverse the flow of fan exhaust air to slow the airplane during the landing roll. Core Cowl The core cowl is an integral component of the thrust reverser. It is constructed in two halves, each linked to the pylon at the top with a latch (required to maintain the fireseal) and held together at 6 o'clock by a latch. The core cowl skin is integral part of the fixed structure skin, and is opened with the reversers. Purpose (2.B.b) The engine cowls provide protection for the engine and accessories, and also ensures smooth airflow around the engine during flight. The CF34-10E nacelle and cowling is protected from lightning and static electricity.

71-00-00 POWERPLANT

Page 14 Aug 05

GE AIRCRAFT ENGINES


CF34-10E

TRAINING MANUAL

71-00-00 POWERPLANT

Page 15 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

AIR MANAGEMENT SYSTEM (AMS) Identification (1.A.a) The Air Management System is located on the left side of the core module and consists of the Nacelle Pressure Regulating and Shutoff Valve (NAPRSOV), High Pressure Regulating and Shutoff Valve (HPRSOV), Low Stage Bleed Check Valve, Pre-cooler, Fan Air Valve and the ducting. Refer to figure, AIR MANAGEMENT SYSTEM. Purpose (2.B.b) The Air Management System provides bleed air for the aircraft environmental control system, aid in starting the opposite engine and for aircraft required anti-ice. Operation (3.C.c) Fifth stage and ninth stage bleed air is utilized to meet the requirements for various bleed configurations. This pressure is sensed in the bleed ducting aft of the firewall. The AMS system continuously monitors engine bleed pressure and will use the 9 th stage bleed air for supply up to an engine condition where the 9 th stage will close down and 5th stage air will become the primary source. Bleed pressures from the low stage supply (5 th stage) is insufficient at lower engine speeds. Bleed pressure from the high stage supply (9 th stage) is too high at higher engine speeds. The bleed pressure is controlled to 45 psig up to 25,000 ft and 35 psig above 25,000 ft using the HPRSOV and the NAPRSOV. The Low Stage Bleed Check Valve is installed in the 5 th stage line to prevent reverse flow of the 9 th stage air into the 5 th stage ducting.


Air is supplied through the pre-cooler to the ECS system and is temperature controlled by the AMS by adjusting the Fan Air Valve which controls the amount of cold fan air flowing across the pre-cooler. A temperature of 204 deg C (400 deg F) is controlled under usual conditions. This temperature is increased during operation of the wing anti-ice to 231 deg C (448 deg F). If the 5th stage air temperature and pressure are insufficient for anti-ice, the AMS will bleed in hotter, higher pressure 9 th stage air to meet the requirements, using the Fan Air Valve for temperature control and the NAPRSOV for pressure control. The FADEC receives bleed configuration discretes via the ARINC bus from the AMS system to determine the current bleed configuration of the aircraft. This is used to vary the N1 setting (thrust) in order to maintain engine ITT within the limits established for the rating.

71-36-00 POWERPLANT

Page 16 June 08

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Fan Air Valve Pre-cooler

Combined Manifold to Precooler

Bellows 9TH Stage Pressure Line

NAPRSOV

HPRSOV

Nacelle Anti-ice

5TH Stage Pressure Line

Low-Stage Bleed Check Valve

AIR MANAGEMENT SYSTEM EFFECTIVITY ALL GE PROPRIETARY INFORMATION

71-36-00 POWERPLANT

Page 17 June 08

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

THERMAL ANTI-ICE SYSTEM (TAI) Identification (1.A.a) The Thermal Anti-ice System is located on left side of the fan case and consists of the following components: 5 th stage air supply duct Thermal Anti-ice (TAI) valve, TAI line pressure transducer, dedicated muscle line from the starter duct for TAI valve operation and a triple swirl nozzle. Refer to figure, THERMAL ANTI-ICE SYSTEM Purpose (2.B.b) The Thermal Anti-ice System supplies hot air from the engine to the inlet cowl to prevent the hazardous formation of ice on the inlet lip.

At the inlet forward bulkhead, the air passes through a triple swirl nozzle and impinges on the inner surface of the inlet lip skin. The air then exits through an exhaust port near the bottom of the inlet. A pressure transducer is mounted in the ducting downstream of the valve to indicate an air supply pressure problem. If the duct pressure falls below a set value, as with a valve failure or duct rupture, a message is sent to notify the crew of a loss of anti-icing. This is a CAS message. If pressure increases beyond a set value, as in a valve failing full open, a message is sent the maintenance record for corrective action. This is not a CAS message.

Operation (3.C.c) The TAI system takes hot air from the 5 th stage bleed ducting at the AMS system. This port is located upstream of the low pressure check valve to make sure that an air supply is always available when the engine is running. On the overhead panel, Engine 1 and 2 push-button switches are used to select or close the correspondent EAI system. The Mode switch will select the mode the anti-ice system will operate. When in AUTO, the engine system will be activated by signals from any of the aircraft ice detectors. When ON is selected, the system will be activated independent of the ice detection system. The valve is spring loaded to the open position so that the system defaults open if there is a loss of control signal or there is not sufficient muscle air pressure from the starter duct to hold the valve closed.


71-30-00 POWERPLANT

Page 18 Sept 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Starter Duct

Air Starter Valve

5TH Stage Air Supply

Muscle Air Line

Pressure Transducer Anti-ice Duct Ice Protection Overhead Panel Air to Triple Swirl Nozzle

Anti-ice Valve

THERMAL ANTI-ICE SYSTEM EFFECTIVITY ALL GE PROPRIETARY INFORMATION

71-30-00 POWERPLANT

Page 19 Sept 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE FIRE ZONES Identification (1.A.a) There are two fire zones for the CF34-10E engine. One is the fan compartment and the other is the core compartment. Refer to figure, ENGINE FIRE ZONES Purpose (2.B.b) The fire zone isolate the fan and core compartments with independent fire detection and fire extinguishing capability.


71-26-00 POWERPLANT

Page 20 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Core Zone Fan Zone

ENGINE FIRE ZONES EFFECTIVITY ALL GE PROPRIETARY INFORMATION

71-26-00 POWERPLANT

Page 21 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FIRE DETECTION LOOPS Identification (1.A.a) The CF34-10E has independent fire loops for the fan fire zone and for the core fire zone. Refer to figure, FIRE DETECTION LOOPS Purpose (2.B.b) The fire detection loops are designed to provide necessary fire detection for the fan and core zones where fires are most likely to occur. Operation (3.C.c) The fire detection loops consist of a pneumatic sensing device which provides for detection due to gas expansion and electrical switch closure (DRL Responders). Two loops within each zone provide redundancy in the event of one loop failure.


71-26-00 POWERPLANT

Page 22 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

DRL Responders

Left Hand Fire Loop

Top View Core Compartment

Fan Compartment

FIRE DETECTION LOOPS EFFECTIVITY ALL GE PROPRIETARY INFORMATION

71-26-00 POWERPLANT

Page 23 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE DRAINS Identification (1.A.a) The engine drain system has lines that collect and transmit waste fluid overboard from accessories, pylon drain cavities, and various systems along the engine. The drain system has two outputs: - Forward drain mast - Aft drain mast Refer to figure, ENGINE DRAINS. The forward mast contains the accessory gearbox component pad drains as well as the oil scupper and pylon drains. Seven tubes are routed to the forward drain mast in the fan zone. Four stainless steel 3/8 inch OD tubes are used to drain the accessory pads of the hydraulic pump, starter, IDG, and fuel pump. A fifth stainless steel 3/8 inch OD tube is used to drain the overfill scupper on the oil tank. The final two drains provide for the hydraulic case drain and any accumulated fluids in the forward pylon through the pylon floor drain. 6 1 0 1 5 0 0 0 0 0 1 7 D S

The masts extend beyond the boundary layer of the fan cowl in order to prevent nacelle streaking with drained fluids. A drip lip is incorporated into the drain mast to prevent the drained fluids from running back on the mast itself. Purpose (2.B.b) The engine drain system lets oil, fuel, hydraulic fluid, water, and vapor to flow overboard through the nacelle structure and makes sure the fluids do not touch the hot engine areas.

The aft mast primarily contains the fuel drains from the core mounted components. Seven 3/8 inch OD tubes are routed to the aft drain mast in the fan zone. Four of the seven stainless steel tubes are used to drain the VBV’s, VSV’s, HPTACC valve, LPTACC valve, and TBV fuel actuators. The remaining tubes are used to drain any accumulated fluids from the extension ring, core pylon, and an oil drain from the forward sump.


71-71-00 POWERPLANT

Page 24 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Forward Drain Mast

AFT Drain Mast Left Cowl

FWD

G F A B C D E F

G

E D

A B C

Right Cowl

FWD Drain #

Component

A

Pylon Fuel Drain

B

Oil Tank

C D

E

7 1 0 1 1 0 0 0 0 0 1 7 D S

Fuel Pump Pad

Fluid

Engine Oil Overflow at Oil Tank Scupper Oil from AGB Carbon Seal or Fuel from Pump Shaft Leakage

Low Point in Pylon, any Fluid Oil from AGB Carbon Seal or Hydraulic Fluid Hydraulic Pump Pad from Pump Shaft Leakage

Component

A

Forward Sump Drain

Engine oil from forward sump seals

B

Pylon Drain

Low point in engine pylon, any fluid

C

VSV/VBV Left Hand Fuel from internal Actuators shaft packing leakage

Pylon Floor Drain

F

IDG Pad

G

Starter Pad

EFFECTIVITY

Oil from AGB Carbon Seal or IDG Oil from Pump Shaft Leakage Oil from AGB Carbon Seal or Starter Oil from Pump Shaft Leakage

Fluid

Drain #

D

TBV Valve

Fuel from Internal Leakage

Not Used E

HPTACC and LPTACC Fuel from internal Valves valve leakage

F

VSV/VBV Right Hand Fuel from internal Actuators shaft packing leakage

G

ALL

Fan Fase Drain

Low point in the extention ring, any fluid

71-71-00

GE PROPRIETARY INFORMATION

ENGINE DRAINS

POWERPLANT

Page 25 June 09

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE MOUNTS Identification (1.A.a) There is a forward and an aft engine mount. The forward engine mount attaches the fan Outlet Guide Vane (OGV) frame to the pylon. The aft engine mount attaches to the engine turbine rear frame and the front frame through a thrust link to the pylon. The engine mounts are designed to allow for thermal expansion of the engine. The mounts also allow for simple engine removal and installation. Refer to figure, ENGINE MOUNTS. Purpose (2.B.b) The engine mounts attach the engine to the pylon and transmit thrust to the aircraft.

8 1 0 1 5 0 0 0 0 0 1 7 D S


71-20-00 POWERPLANT

Page 26 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Aft Mount Forward Mount Yoke Thrust Link Forward Mount

Shear Links

9 1 0 1 1 0 0 0 0 0 1 7 D S

Thrust Yoke

ENGINE MOUNTS EFFECTIVITY ALL GE PROPRIETARY INFORMATION

71-20-00 POWERPLANT

Page 27 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FAULT DETECTION STRATEGY


00-00-00 FAULT DETECTION STRATEGY

Page 1 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FAULT DETECTION STRATEGY Objectives Given an objective exercise, the student will identify: - The FADEC fault detection system (1.A.a) - The dispatch level (1.A.a) - The built-in test equipment (BITE) (1.A.a) - The recent faults (1.A.a) - The fault history (1.A.a) - The identification configuration (1.A.a) - The ground test (1.A.a) - The input monitoring (1.A.a) - The maintenance message breakdown (1.A.a) Given an objective exercise, the student will select the operation of: - FADEC fault detection system (3.C.c)

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Page 2 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL




Page 3 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FADEC FAULT DETECTION SYSTEM Identification (1.A.a) The FADEC will monitor inputs from various sources and perform tests for failures and overlimit conditions. The FADEC will also monitor data to determine when to request trend data saves, when limit exceedances are to be tracked and reported. The FADEC will also work in conjunction with the MAU, CMC, MFD and the CCD to provide engine test capability for maintenance personnel. The FADEC will supply data to other aircraft avionics via ARINC 429. This data will aid in detection, reporting and storage of faults, engine dispatch level and exceedance information. Data for trend data records will also be supplied to the aircraft. Data will be provided for the Flight Data Recorder (FDR).

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Operation (3.C.c) The FADEC will perform diagnostics on its input data to determine if any failures exist. If a failure exists, and if the failure is one that the pilot needs to be aware of, a message or indication will be displayed to the pilot on the EICAS. If a failure condition is not severe and does not directly affect engine operation, the engine dispatch level is calculated and a message will be transmitted for display on the appropriate device (EICAS or MFD). Some engine related EICAS messages or indications and maintenance faults require the MAU or the SPDA to perform the checks as the required information is not


available to the FADEC. Ignition CAS Icons will be displayed to the pilot on the EICAS display if all the given criteria have been met. Fault data will be saved in the CMC for display on the CMC-driven MFD display. The CMC will display fault related information to maintenance personnel on the MFD via the CMC. The fault messages will correspond to the EICAS messages and are to aid engine fault troubleshooting and correction. Fault codes for all detected faults (up to 16 per channel) will be displayed on an MFD-driven MFD display. The purpose of this display is to provide a simple backup to the CMC. If the CMC is not available, the MFD fault display will allow maintenance personnel to determine what engine faults are active and to allow reference of troubleshooting procedures for the fault.


Page 4 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

Aircraft

TRAINING MANUAL

Engine

MFD A EICAS

R I N C

MAU 4 2 CMC 5 0 0 1 1 0 0 0 0 0 0 0 D S

9

FADEC FAULT DETECTION SYSTEM



Page 5 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

DISPATCH LEVEL Identification (1.A.a) The FADEC will determine the health or failure of the items that it can monitor. The FADEC does not monitor several of the engine-related components ; the MAUs and SPDAs are responsible for their monitoring. Some engine failures have dedicated EICAS messages or indications. Those failures must be dispositioned by referring to the MMEL. The FADEC logic will not set the dispatch bits for these indications.

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For those engine failures that are monitored by the FADEC and where no other cockpit indications are set, the FADEC will use the maintenance fault words to set dispatch limitation bits. The FADEC will scan the maintenance fault words and search to see if any local channel faults bits are set. If any are set then fault bits from the local channel, cross channel and a set of masks are used to determine the correct dispatch limitation level. After the dispatch level is determined, the appropriate flag will be set. These flags will be used to set bits in the ARINC 429 output status words to set the appropriate cockpit indication of dispatch limitation level. Dispatch Level Indication Latching New dispatch levels will not be latched for NVM storage when the engine is sub idle. If the engine is sub idle, if a fault is seen, the dispatch bit will be set if not already latched in NVM. Also when sub idle, if the fault clears, the


respective dispatch bit will also clear unless it is already latched in NVM. If a fault that the FADEC is aware of exists when the engine reaches idle for 10 seconds, that fault will cause a dispatch bit to be latched and stored in NVM. Dispatch Level Indication Clearing The dispatch level indication bits will be cleared when maintenance personnel request the maintenance faults to be cleared through the MFD, CCD user interface. Dispatch Fault Levels There are 5 defined dispatch fault levels: In-Flight Warning A - No Dispatch B - Short Time C - Long Time D - Economic The In-Flight Warning indicates to the flight crew a condition that requires pilot action or notification. The No Dispatch level indicates that maintenance action is required before the aircraft can perform a flight. This message is displayed on the EICAS display.


Page 6 June 08

GE AIRCRAFT ENGINES

Fault Code

CF34-10E

Dispatch Level

EICAS

VSV DUAL CMDCURRENT DISAG

IN FLIGHT WARNING

ENGINE X CONTROL FAULT

ECP

ECPFAULT

NO DISPATCH

Ex NO DISPATCH

TBV

TBV SENSOR OUT OF RANGE

SHORT DISPATCH

Ex SHORT DISPATCH

Local channel TBV feedback sensor has failed. Cross channel sensor is used.

T2 OUT OF RANGE

LONG DISPATCH

(NONE)

The local channel engine T2 sensor is out of range of -100 to 199 F.

LPT CMD-CURRENT DISAGREE

ECONOMIC

(NONE)

Single channel failure in the torque motor driver circuit.

LRU

7321 5416 Exz

7321 7526 Exz

7321 5521 Exz

7321 6722 Exz

T12

7321 5324 Exz LPTACC

7 0 0 1 1 0 0 0 0 0 0 0 D S

TRAINING MANUAL

CMC Fault Message

Description Dual channel failure in the torque motor drive circuits Data from both channels ECP are missing or invalid.

X=ENGINE NUMBER Z=FADEC CHANNEL

DISPATCH LEVELS



Page 7 Feb 09

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

The Short Time dispatch level indicates that at least one short time dispatch fault exists and must be noted in the logbook by the pilot. This fault is allowed to persist for 10 days or 150 flight hours before it must be corrected. This message is displayed on EICAS. Certain short time faults, when combined, can set the No Dispatch level. These are known as “Alpha” faults. The Long Time dispatch level indicates that at least one long time dispatch fault exists and the message is displayed on the MFD which is not normally viewed by the pilot.The means for checking long time dispatch faults will be to interrogate the MFD-driven Engine Maintenance Page at "A" check. The Economic time dispatch level indicates that at least one economic dispatch fault exists. This dispatch level does not require repair and is included as failure in items that effect economic aspects. The means for checking economic faults is to interrogate the MFD-driven Engine Maintenance Page at "A" check. 8 0 0 1 5 0 0 0 0 0 0 0 D S

Final determination of Short Time and Long Time dispatch is done by pilot or maintenance crews referencing to the MMEL for a dispatch or no dispatch condition.



Page 8 June 08

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL




Page 9 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Multi Functional Display Engine Maintenance Page

Identification (1.A.a) Engine fault data will be saved in the CMC for display on the CMC-driven MFD display. The CMC will display fault related information to maintenance personnel on the MFD via the CMC. The fault messages will correspond to the EICAS messages and are to aid engine fault troubleshooting and correction. The engine maintenance page will display the following items: - FADEC detected dispatch limitations - Exceedance events - Fault codes saved in the CMC for engine/ FADEC



Page 16 Mar 06

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

MUTI FUNCTIONAL DISPLAY – ENGINE MAINTENANCE PAGE



Page 17 Mar 06

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

BUILT-IN TEST EQUIPMENT (BITE) Identification (1.A.a) The FADEC has three BIT modes designed to test the computer functions of the main CPU and the Autonomous Input Processor (AIP) as well as the output drivers and ARINC transceiver chip. - Power On BIT (POBIT) - Ground Start BIT (GSBIT) - Normal Periodic BIT (NPBIT) POBIT POBIT is initiated automatically after any FADEC power on reset. This test is designed to test the most critical computer functions while allowing the FADEC to quickly come on line. POBIT takes a maximum of 0.246 seconds to execute. GSBIT GSBIT is executed on the ground if the engine is not running when any of these conditions exist: - After the completion of POBIT. - When the engine is being shutdown. 0 1 0 1 5 0 0 0 0 0 0 0 D S

GSBIT takes a maximum of 1.982 seconds to execute. NPBIT NPBIT is run continuously in foreground or background mode. This test is designed to continuously monitor the health of the processors and the computer.



Page 10 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL




Page 11 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

3 1 0 1 1 0 0 0 0 0 0 0 D S



Page 13 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

If the CMC was not available during the flight and is plugged in and is operational after a flight, the CMC will report the peak and duration information of any exceedances detected by the FADEC, but no other data is available.

4 1 0 1 5 0 0 0 0 0 0 0 D S



Page 14 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Maintenance Page MAINTENANCE MESSAGE DETAIL 35 CHARACTERS FAULT NAME: WASTE SERV PNL SW/WWSC/MAU3 FAULT FAULT TYPE: PROBE / SENSOR INTERNAL/INTERFACE/PROBE/SENSOR FAULT CODE: 38325784PNL FAULT CODE LRU: HIGHEST TO LOWEST PROBABILITY LRU (S) AT FAULT: VWS Service Panel Switch Water & W aste System Controller MAU3 – Generic I/O Module (slot 10) Aircraft Wiring

SYMPTOM DETAILS, NOT REQUIRED SYMPTOM: Check MAU3 for fault reporting. Service panel door switch may be improperly adjusted or malfunctioning.

DOCUMENTS

LINK TO FAULT ISOLATION MANUAL

MAINTENANCE MESSAGE OCCURRENCES: 5 1 0 1 1 0 0 0 0 0 0 0 D S

ACTIVE INACTIVE ACTIVE MAIN MENU

MAY 13, 2000 19:20:02 LEG:1 CRUISE MAY 13, 2000 19:10:33 LEG:1 CLIMB MAY 13, 2000 19:09:05 LEG:1 TO RUN PREV

CENTRAL MAINTENANCE COMPUTER (CMC) EFFECTIVITY ALL GE PROPRIETARY INFORMATION


Page 15 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FAULT HISTORY Identification (1.A.a) The CMC will record trend points and these points will not be cleared unless overwritten by newer trend points. The CMC will save up to 200 trend points total. 200 trend points will allow for a weekly data download interval assuming a maximum of 10 flights a day, 3 trend points a flight (2 takeoff, 1 cruise), 7 day operation. The CMC has NVM to save engine-related data indefinitely or until cleared or overwritten. The CMC will save a set amount of data for each fault. The fault history can be accessed through the CMC by selecting the Fault History by Date or Fault History by ATA. Selecting the following soft key for Historical by Date will then access a Month selection (three month span), then by a Date, and then to Flight Leg. It is then broken down to Warnings, Cautions, or Advisorys for that flight leg. Selecting the following soft keys for Historical by ATA will then access an ATA list and then to a fault list related to that ATA. 6 1 0 1 5 0 0 0 0 0 0 0 D S

- Air/Ground Status - Fault Code Snapshot data storage in the FADEC NVM only, the CMC does not store engine data for engine maintenance faults. Exceedance Data The CMS will record the following data for each event: - Aircraft Serial Number - Date - Time - FADEC location ID (SDI code) - Flight Leg - Flight Phase Data, as monitored and configured in the CMC. This data may vary based on function (exceedance type or trend data point).

Maintenance Faults The standard CMC fault storage function will record the following data for each fault: - Aircraft Serial Number - Date - Time - FADEC location ID (SDI code)



Page 18 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

7 1 0 1 1 0 0 0 0 0 0 0 D S



Page 19 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL




Page 23 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE GENERAL


72-00-00 ENGINE GENERAL

Page 1 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE GENERAL Objectives Given an objective exercise, the student will identify: - The sumps and frames (1.A.a) - The engine bearings (1.A.a) - The engine aerodynamic stations (1.A.a) - The fan module and BSI ports (1.A.a) - The fan blades and spinners (1.A.a) - The accessory gearbox and component location (1.A.a) - The core module and BSI ports (1.A.a) - The LPT module and BSI ports (1.A.a) Given an objective exercise, the student will select the purpose of: - The sumps and frames (2.B.b) - The engine bearings (2.B.b) - The fan module and BSI ports (2.B.b) - The fan blades and spinners (2.B.b) - The accessory gearbox and component location (2.B.b) - The core module and BSI ports (2.B.b) - The LPT major module and BSI ports (2.B.b)

2 0 0 1 5 0 0 0 0 0 2 7 D S



Page 2 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL




Page 3 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

SUMPS AND FRAMES Identification (1.A.a) The sumps are part of the engine oil system and the engine vent system. The engine has two sumps: - Forward sump - Rear sump The forward sump is part of the fan frame. The rear sump is part of the turbine rear frame Purpose (2.B.b) The forward sump supports the fan and booster rotor assemblies and connects the accessory gearbox through the inlet gearbox. The sump is internal to the fan frame assembly which provides the main forward support for mounting the engine to the aircraft through the forward engine mounts. Bearings No.1, No.2, No, 3B and No.3R are in the forward sump.

4 0 0 1 5 0 0 0 0 0 2 7 D S

The rear sump supports the HPT and LPT rotors through the turbine rear frame which provides the main rear support for mounting the engine to the aircraft through the rear engine mounts. Bearings No.4 and No.5 are in the rear sump.



Page 4 May 07

GE AIRCRAFT ENGINES

Radial Drive Shaft TAI Duct

CF34-10E

TRAINING MANUAL

ALF

5 0 0 1 1 0 0 0 0 0 2 7 D S

To Drain Mast



Page 5 May 07

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL




Page 6 Mar 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FLA

Aft Sump Oil Supply Tube

7 0 0 1 1 0 0 0 0 0 2 7 D S

Oil Supply



Page 7 Sept 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE BEARINGS Identification (1.A.a) The engine has six main bearings (two ball and four roller) in the sumps. Bearings No.1B, No.2R, and No.3B/ R are in the forward sump. Bearings No.4R and No.5R are in the rear sump. Purpose (2.B.b) Bearings absorb the axial and radial loads from the N1 and N2 shafts. The No.1 ball bearing and No.2 roller bearing support the front of the fan shaft. The No.3 ball bearing and No.3 roller bearing support the forward end of the HPC shaft in the inlet gearbox. The No.4 roller bearing supports the HPT rear shaft. The No.5 roller bearing supports the end of the LPT shaft. Training Information Points (3.E.e) The main engine bearings are made of M50 material.

8 0 0 1 5 0 0 0 0 0 2 7 D S



Page 8 Jan 08

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

9 0 0 1 1 0 0 0 0 0 2 7 D S



Page 9 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE AERODYNAMIC STATIONS Identification (1.A.a) Engine aerodynamic stations are at axial locations. They follow a Station Designation System which is used to easily identifying certain parameters like temperatures and pressures.

0 1 0 1 5 0 0 0 0 0 2 7 D S



Page 10 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

LPTACC VALVE

1 1 0 1 1 0 0 0 0 0 2 7 D S



Page 11 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Engine Modules Identification (1.A.a) The engine modules consist of the following: - Fan module module assemb assembly ly - Core module module assemb assembly ly - LPT module module assem assembly bly - Accessory Accessory gearbox gearbox assembly

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Page 12 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

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Page 13 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FAN MODULE AND BSI PORTS Identification (1.A.a) The fan module is in the front of the engine and is made of the fan stator, the fan, and the Low Pressure Compressor (LPC) and Inlet Gearbox. The LPC borescope inspection location is identified as port S0. Port S0 is found aft of the fan blades in the secondary air area at approximately 3 o’clock and 9 o’clock position, between the platforms of an Outlet Guide Vane (OGV). There is no borescope plug at this location. Borescope port S0 is used to see: - Stage 2 vane vane trailing edge - Stage Stage 3 vane vane - Stage 3 rotor blade blade trailing edge edge - Stage 4 vane vane leading leading edge - Stage 4 rotor blade blade leading leading edge

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Purpose (2.B.b) The fan increases the speed of the intake air. A splitter fairing divides the air into the primary and secondary airflows. The primary air enters the booster compressor to provide pressurized air to the High Pressure Compressor (HPC). The secondary air is accelerated through the core cowl to atmosphere. The Inlet gearbox provides rotation from the compressor rotor to the transfer gearbox through a radial drive shaft. Borescope inspection port S0 is for the inspection of the booster compressor.



Page 14 March 06

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Inlet Gearbox

Low Pressure Compressor

o

Fan Blades

Fan Stator Case

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S0 Borescope 3:30 Splitter Fairing

FAN MODULE



Page 15 March 06

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FAN BLADES AND SPINNER Identification (1.A.a) The fan consists of a 53" diameter single stage fan rotor made up of 24 wide chord blades. The blades are held in position by a double bore fan disk bolted to a fan forward shaft splined to the LPT shaft. Covering the fan disk and attached to it is a one piece spinner with 36 balance weights threaded circumferentially into captive nuts. Purpose (2.B.b) The fan blades are components of the fan module which compresses primary air and pushes secondary airflow around the engine. The spinner helps set the ideal airflow path into the engine inlet and is specially shaped to deflect ice and FOD from the core inlet. The spinner is a “coniptical” design which is a compromise between a conical, used on the CFM56-2, and the elliptical, used on the CF6.

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Training Information Points (3.E.c) With any procedure that requires the removal of the fan blades, ensure that they are reinstalled back in the same position to maintain fan rotor balance. For any procedure that requires removal of the inlet components, check the condition of the blade lubrication before reinstallation. Loss of lubrication can lead to fan vibration.



Page 16 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

RELUBE BLADES AT 1500-3000 CYCLES

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Page 17 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ACCESSORY GEARBOX Identification (1.A.a) The Accessory Gearbox (AGB) is on the left side of the engine fan case. The following components are at the front of the AGB: - Permanent Magnet Alternator (PMA) - Engine air starter - Hydraulic pump - Integrated Drive Generator (IDG) - N2 speed sensor (located on top of AGB) The following components are at the aft of the AGB: - Lube filter module - Fuel pump - Lube and scavenge pump - N2 cranking pad Purpose (2.B.b) Power to drive all AGB accessories is supplied from the HPC rotor by means of an inlet gearbox (IGB) through a radial drive shaft to a transfer gearbox (TGB) then through a horizontal drive shaft to the AGB. 8 1 0 1 5 0 0 0 0 0 2 7 D S



Page 18 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

AGB FRONT VIEW

AGB REAR VIEW

TRAINING MANUAL

DRIVE SYSTEM

N2 Sensor

Inlet Gearbox (forward sump) Radial Drive Shaft

Hydraulic Pump Axis F Fuel Pump Axis E N2 Cranking Pad

Starter Axis D

Input Drive Axis C 9 1 0 1 1 0 0 0 0 0 2 7 D S

Horizontal Drive Shaft

IDG Axis G Lube/Scavenge Pump Axis J

Alternator Axis J

Transfer Gearbox

ACCESSORY GEARBOX AND COMPONENT LOCATION EFFECTIVITY ALL GE PROPRIETARY INFORMATION


Page 19 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

CORE MODULE AND BSI PORTS Identification (1.A.a) The core module consists of the following: - The high pressure compressor (HPC) - The combustor - The high pressure turbine (HPT)

Purpose (2.B.b) The HPC increases the pressure of the air from the LPC and sends it to the combustor. The HPC also supplies bleed air for the aircraft pneumatic system and the engine air system.

There are 16 borescope ports on the core module.

The combustor mixes air from the compressor and fuel from the fuel nozzles. This mixture of air and fuel is burned in the combustion chamber to make hot gases. The hot gases go to the HPT.

The borescope ports on the HPC are identified as follows: - S1 through S9 are located along the right side at the 2:00 position. The borescope ports for inspection of the combustion chamber are identified as follows: - S10 (2 o’clock) Combustion Liner - S11 (5:30 position) Combustion Liner - S12 (10 o’clock) Combustion Liner - S13 (Right igniter plug port) (4 o’clock) - S14 (Left igniter plug port) (8 o’clock)

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The HPT changes the energy of the hot gases into mechanical energy. The HPT uses the mechanical energy to turn the HPC rotor and the accessory drive. Borescope ports on the core major module are for inspection of the compressor rotor, combustion chamber, and the HPT.

The borescope ports for inspection of the HPT section/LPT stage 1 nozzle are identified as follows: - S15 (3 o’clock) - S16 (9 o’clock))



Page 20 May 07

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

S15 S16 S10 S11 S12 S1

S2

S3

S4

S5

S6

S7 S8

S13 S14

S9 LPT Nozzle HPT Nozzle

HPC

Combustion Liner

HPT Rotor

CORE MODULE AND BSI PORTS EFFECTIVITY ALL GE PROPRIETARY INFORMATION


Page 21 March 06

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

LPT MODULE AND BSI PORTS Identification (1.A.a) The low pressure turbine (LPT) module is the aft of the core module. The LPT is a four-stage turbine. There are two borescope ports on the LPT module. The borescope ports used to inspect the LPT are identified as follows: - S17 (3 o’clock) - S18 (3 o’clock) Purpose (2.B.b) The LPT changes the energy of hot gases into mechanical energy. The LPT uses the mechanical energy to turn the fan and the LPC rotor. The LPT components seen through borescope ports S17 andu S18 are as follows: - S17 Stage 1 LPT rotor blade trailing edge - S17 Stage 2 LPT nozzle segments - S17 Stage 2 LPT rotor blade leading edge - S18 Stage 2 LPT rotor blade trailing edge - S18 Stage 3 LPT nozzle segments - S18 Stage 3 LPT rotor blade leading edge

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Page 22 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

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Page 23 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FUEL AND CONTROL

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 1 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FUEL AND CONTROL

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Objectives Given an objective exercise, the student will identify: - The fuel and control (1.A.a) - The location of the main fuel pump (1.A.a) - The location of the main fuel filter (1.A.a) - The location of the filter delta pressure switch (1.A.a) - The location of the main fuel/oil heat exchanger (1.A.a) - The location of the servo fuel/oil heat exchanger (1.A.a) - The location of the integrated drive generator - fuel/oil HX (1.A.a) - The location of the fuel metering unit (1.A.a) - The location of the fuel manifold (1.A.a) - The location of the fuel injectors (1.A.a) - The engine control subsystem (1.A.a) - The engine control interfaces (1.A.a) - The typical engine control loop (1.A.a) - The location of thrust lever (1.A.a) - The auto throttle interface (1.A.a) - Selected components of the engine fuel control system (1.A.a) - The location of the FADEC (1.A.a) - The location of the T12 sensor (1.A.a) - The location of the T25 sensor (1.A.a) - The location of the T3 sensor (1.A.a) - The location of the delta T3 sensor (1.A.a) - The location of the Tcase sensor (1.A.a) - The location of the fan speed sensor - N1 (1.A.a) - The location of the core speed sensor - N2 (1.A.a) - The location of the permanent magnet alternator (PMA) (1.A.a) - The location of the engine configuration plug (1.A.a)

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 2 Aug 05

GE AIRCRAFT ENGINES

3 0 0 1 5 0 0 0 0 0 3 7 D S

CF34-10E

TRAINING MANUAL

Given an objective exercise, the student will select the purpose of: - The fuel and control (2.B.b) - The main fuel pump (2.B.b) - The main fuel filter (2.B.b) - The filter delta pressure switch (2.B.b) - The main fuel/oil heat exchanger (2.B.b) - The servo fuel/oil heat exchanger (2.B.b) - The integrated drive generator - fuel/oil HX (2.B.b) - The fuel metering unit (2.B.b) - The fuel manifold (2.B.b) - The fuel injectors (2.B.b) - The engine control subsystem (2.B.b) - The engine control interfaces (2.B.b) - The thrust lever (2.B.b) - The auto throttle interface (2.B.b) - Selected components of the engine fuel control system (2.B.b) - The FADEC (2.B.b) - The T12 sensor (2.B.b) - The T25 sensor (2.B.b) - The T3 sensor (2.B.b) - The delta T3 sensor (2.B.b) - The Tcase sensor (2.B.b) -The fan speed sensor - N1 (2.B.b) - The core speed sensor - N1 (2.B.b) - The permanent magnet alternator (PMA) (2.B.b) Given an objective exercise, the student will select the operation of the: - The fuel and control (3.C.c) - The main fuel pump (3.C.c) - The filter delta pressure switch (3.C.c) - The fuel metering unit (3.C.c) - The fuel injectors (3.C.c)

EFFECTIVITY ALL

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Page 3 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

- The auto throttle interface (3.C.c) - Selected components of the engine fuel control system (3.C.c) - The FADEC (3.C.c)

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EFFECTIVITY ALL

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FUEL AND CONTROL

Page 4 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL


EFFECTIVITY ALL

73-00-00


FUEL AND CONTROL

Page 5 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FUEL AND CONTROL Identification (1.A.a) The CF34-10E engine fuel and control system supplies fuel for all engine thrust and control operations. All fuel and control components are on the engine. The fuel and control system is divided into these three subsystems: - Engine fuel distribution system - Engine control subsystem - Engine control system Refer to figure, FUEL AND CONTROL.

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Engine Fuel Distribution System The engine fuel system consists of the following components: - Main fuel pump - Main fuel filter - Filter delta pressure switch - Main fuel/oil heat exchanger - Servo fuel/oil heat exchanger - Integrated drive generator - fuel/oil HX - Fuel metering unit (FMU) - Fuel manifolds - Fuel injectors Engine Control Sub-System Engine control subsystem consists of the following: - Engine control interfaces

EFFECTIVITY

- Thrust lever - Auto throttle interface Engine Control System The CF34-10E engine control system is a computer based electronic engine control system. It is composed of a two channel Full Authority Digital Engine Control (FADEC), a Fuel Metering Unit (FMU), a Permanent Magnet Alternator (PMA), engine sensors, Variable Stator Vane (VSV) actuator, Variable Bleed Valve (VBV) actuator, Transient Bleed Valve (TBV) actuator, High Pressure Turbine Clearance Control Valve (HPTCC) Actuator, Low Pressure Turbine Clearance Control Valve (LPTACC) and an ignition system for each engine. The system controls the engine in response to thrust command inputs from the aircraft and provides information to the aircraft for cockpit indication, maintenance reporting and engine condition monitoring. Purpose (2.B.b) The purpose of the engine fuel system is to provide scheduled fuel to the engine to provide combustion required for propulsion power. The engine fuel system includes the management of the fuel provided by the engine control system and the delivery of the aircraft supplied fuel. The delivery system provides pressurization, heating, and filtering of the fuel, and ultimate delivery into the combustion chamber for burning.

ALL

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FUEL AND CONTROL

Page 6 May 07

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

7 0 0 1 1 0 0 0 0 0 3 7 D S

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 7 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Main Fuel Pump 9 0 0 1 1 0 0 0 0 0 3 7 D S

Motive Return flow Motive Element

Inlet flow

Boost Element

Main Element

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 9 Aug 06

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

MAIN FUEL PUMP Identification (1.A.a) The main fuel pump is mounted on the left side of the engine (outboard) aft face of the accessory gearbox. The bearings in the pump are all lubricated by the fuel and do not rely on an outside source of oil. During windmill operation, the bearings remain lubricated by the recirculating fuel in the system. The FMU unloads the pump to a lower discharge pressure during wind milling to reduce the heat load in the system and pump bearing loads.

remainder leaving the pump and being sent through the fuel/oil heat exchanger. The fuel entering the secondary gear stage is further pressurized before leaving the pump to be sent to the aircraft as motive flow. The fuel leaving the fuel/oil heat exchanger then reenters the fuel pump and is turns to the pump and is further pressurized by the high-pressure gear stage. The fuel then flows out again where it is routed through the fuel filter then routed to the inlet of the FMU.

Refer to figure, MAIN FUEL PUMP. Purpose (2.B.b) The main fuel pump is designed to provide sufficient fuel flow and pressure to meet engine burn flow requirements as well as flow for engine air system actuators through the FMU. The fuel pump also, provides motive flow fuel back to the aircraft for ejector operation.

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Operation (3.C.c) Fuel from the aircraft enters the centrifugal stage of the fuel pump. In the fuel pump, the pressure of the fuel is boosted to provide adequate filling of the downstream gear stages. The flow is then split, with some fuel going on to the secondary high pressure gear stage and the

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 10 May 07

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

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EFFECTIVITY ALL

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FUEL AND CONTROL

Page 11 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

MAIN FUEL FILTER Identification (1.A.a) The main fuel filter is located downstream of the main engine pump. The filter is rated at 10 microns nominal (30 microns absolute). The filter contains a self relieving feature in the event it becomes blocked it will not prohibit flow to the metering valve and combustor. The filter element has been sized to ensure adequate protection of the fuel system components for the duration specified in the engine maintenance manual. The filter and fuel system components have been tested with contaminated fuel to verify their capability to operate with the worst case contaminatation levels. The filter contains a differential pressure switch that provides indication that the filter is reaching its contamination capacity. It’s set point is set to activate before the filter bypass valve is activated. The filter is located in such a way as to be accessible for filter element replacement. The filter bowl contains a drain plug that can be removed to drain the filter bowl before removing the filter element. Refer to figure, MAIN FUEL FILTER. 2 1 0 1 5 0 4 0 1 1 3 7 D S

Purpose (2.B.b) All main pump flow passes through the main fuel filter before entering the FMU both to the servo heat exchanger and metering valve supply.

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 12 Sept 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

3 1 0 1 1 0 4 0 1 1 3 7 D S

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 13 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Identification (1.A.a) The fuel filter delta pressure switch is on the fuel filter head. The fuel filter delta pressure switch is a discrete switch. Refer to figure, FILTER DELTA PRESSURE SWITCH. Purpose (2.B.b) The fuel filter delta pressure switch sends a signal to the flight deck through the MAU when the differential pressure across the filter reaches a certain value. Operation (3.C.c) The fuel filter delta pressure switch shall be open for normal operation and shall be closed when the pressure drop across the fuel filter exceeds the specified limit to indicate that the filter is at impending bypass.

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EFFECTIVITY ALL

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FUEL AND CONTROL

Page 14 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

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FUEL FILTER IMPENDING BYPASS SENSOR EFFECTIVITY ALL

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FUEL AND CONTROL

Page 15 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

MAIN FUEL/OIL HEAT EXCHANGER Identification (1.A.a) The main fuel/oil heat exchanger is located on the fan case at 2 o’clock, just above the oil tank. Refer to figure, MAIN FUEL/OIL HEAT EXCHANGER. Purpose (2.B.b) The main fuel/oil heat exchanger provides cooling of the engine oil. Heat transfer takes place whenever the core is rotating and providing fuel and oil flow. The fuel used to cool the oil is boost pressure coming from the main fuel pump along with bypass fuel returned from the FMU. The fuel then returns to the inlet of the primary pump element in the main fuel pump.

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EFFECTIVITY ALL

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FUEL AND CONTROL

Page 16 Sept 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

7 1 0 1 1 0 4 0 1 2 9 7 D S

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 17 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

SERVO FUEL/OIL HEAT EXCHANGER Identification (1.A.a) The servo fuel/oil heat exchanger is located on the fan case at 4 o’clock and aft of the oil tank. Refer to figure, SERVO FUEL/OIL HEAT EXCHANGER. Purpose (2.B.b) The servo fuel/oil heat exchanger provides heat, rejected from the engine lubrication oil, to the fuel routed to the servo elements in the Fuel Metering Unit (FMU) for the prevention of icing in the servo system.

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EFFECTIVITY ALL

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FUEL AND CONTROL

Page 18 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

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EFFECTIVITY ALL

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FUEL AND CONTROL

Page 19 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

INTEGRATED DRIVE GENERATOR - FUEL/OIL HX Identification (1.A.a) The Integrated Drive Generator (IDG) fuel/oil heat exchanger is on the fan case at the 5 o'clock position. The IDG fuel/oil heat exchanger is a housing containing separate passages for the IDG oil and fuel flow. Refer the figure, INTEGRATED DRIVE GENERATOR FUEL/OIL HX Purpose (2.B.b) The IDG fuel/oil heat exchanger (in conjunction with the IDG air/oil heat exchanger) maintains IDG oil temperature within an acceptable range.

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EFFECTIVITY ALL

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FUEL AND CONTROL

Page 20 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Fuel in from FMU Bypass

Heat Exchanger

Fuel out to Main Fuel/Oil Heat Exchanger Oil Out to IDG

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Oil in from Air/Oil HX

INTERGRATED DRIVE GENERATOR – FUEL/OIL HEAT EXCHANGER

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 21 Sept 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FUEL METERING UNIT (FMU) Identification (1.A.a) The Fuel Metering Unit (FMU) is isolation mounted onto brackets which are attached to the fan case at approximately the 5 o’clock position, aft looking forward. Refer the figure, FUEL METERING UNIT. Purpose (2.B.b) The primary purpose of the FMU is to provide accurate metered fuel flow to the engine for combustion. The FMU also provides fuel flow for variable stator (VSV) and bleed actuation (VBV), operation of the high pressure turbine clearance control valve and the transient bleed valve, plus provides fuel shutoff for normal engine shutdown and for over speed protection. The metering valve position is used for fuel flow.

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Operation (3.C.c) The pump main gear stage total flow passes through the main fuel barrier filter and servo wash filter. The servo flow coming through the wash filter is directed separately to the FMU where it is used to supply the Electro-Hydraulic Servo Valves (EHSV). The EHSVs operate the fuel metering valve, VSV actuators, VBV actuators, HPTCC valve, and the TBV.

fuel then enters the bypass valve and is either returned to the pump or passes through the valve to metering valve and the combustor. The metering valve is positioned by the FADEC to provide the proper burn flow to combustor nozzles. Metered fuel flow exits the FMU and passes through the fuel manifold, where it is distributed to the 20 fuel injectors. The bypass valve varies the amount of bypass flow accordingly by maintaining a constant 50 psid across the metering valve. A shutoff/over speed valve is located at the discharge of the FMU to shutoff fuel flow to the engine. The valve is used to provide normal shutdown and over speed shutdown of the engine. The FMU includes an over speed solenoid and a shutoff over speed valve that operates in response to a signal from the FADEC.

The main flow discharging from the fuel filter is then directed to the inlet of the FMU. The fuel enters the FMU where it passes through the inlet pressurizing valve. The

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 22 Sept 05

GE AIRCRAFT ENGINES

3 2 0 1 1 0 2 0 1 2 3 7 D S

CF34-10E

TRAINING MANUAL

Channels A&B

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 23 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FUEL MANIFOLDS Identification (1.A.a) The fuel manifolds are located around the circumference of the combustion chamber frame, just aft of the forward flange. Refer to figure, FUEL MANIFOLDS. Purpose (2.B.b) The fuel manifolds carry metered fuel from the FMU to the fuel injectors. Fuel coming from the FMU enters the fuel manifolds, and is distributed to the 20 fuel injectors.

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EFFECTIVITY ALL

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Page 24 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Metered Fuel From FMU

Right Manifold

Left Manifold

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FUEL MANIFOLDS EFFECTIVITY ALL

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Page 25 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FUEL INJECTORS Identification (1.A.a) The 20 fuel injectors are equally spaced around the circumference of the combustion chamber frame. Refer to figure, FUEL INJECTORS. Purpose (2.B.b) The fuel injectors atomize the metered fuel from the FMU into the domed combustor. Operation (3.C.c) Fuel from the FMU is sent to each fuel injector through the fuel manifold. Based upon the fuel pressure delivered, a distributor valve located in the fuel injector ports fuel to the primary and secondary fuel flow circuits. The primary circuit is used during start and low power. The secondary circuit provides additional flow at high power. Each fuel injector also has a check valve that closes at engine shutdown to prevent the manifolds from draining into the combustor.

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EFFECTIVITY ALL

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FUEL AND CONTROL

Page 26 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

7 2 0 1 1 0 6 0 1 1 3 7 D S

EFFECTIVITY ALL

73-11-06


FUEL AND CONTROL

Page 27 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL


EFFECTIVITY ALL

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FUEL AND CONTROL

Page 29 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE CONTROL INTERFACES Identification (1.A.a) The aircraft provides engine thrust and control commands and aircraft flight and status information to the engine control systems as described below: 1. Thrust lever position is provided to each FADEC channel via an electrical RVDT signal 2. The following hardwired discretes are provided to each FADEC channel: - Stop switch signal - Weight on wheels - Engine id - Application id - Ground maintenance override

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3. The following data is provided to each FADEC channel on digital ARINC 429 serial data busses: - Air data (total air temperature, altitude, mach number) - Bleed system configuration discretes - Electronic N1 trim - Takeoff Data Set (TDS) and Thrust Rating Selector (TRS) data - Discrete inputs (start, ignition, wind shear warning, ice detector, fire handle, maintenance requests) - Weight on wheels and wheel speed data - Aircraft monitored propulsion system parameters (oil quantity, oil temperature, oil pressure, N1 vibe, N2 vibe)

EFFECTIVITY

- Aircraft configuration data (landing gears position, flaps position) - Fault status information on aircraft systems Engine condition and parameter status are interfaced to the aircraft as follows: 1. Engine condition and status information is transmitted to the aircraft via ARINC 429 serial data busses. 2. Some engine condition signals from the engine to the cockpit are hardwired to the aircraft. Purpose (2.B.b) The aircraft provides engine thrust and control commands and aircraft flight and status information to the engine control systems. The electronic control system also provides airframe interfaces for engine condition monitoring, fault warning, starter cutout control and environmental conditioning cutout control.

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FUEL AND CONTROL

Page 30 Sept 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

AIRCRAFT DATA TO FADEC ANALOG INPUTS - THRUST LEVER ANGLE - 28 VDC AIRCRAFT POWER DISCRETES TO FADEC - FADEC START/STOP SWITCH - WHEIGHT ON WHEELS - ENGINE ID - APPLICATION ID

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ENGINE CONTROL INTERFACES EFFECTIVITY ALL

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FUEL AND CONTROL

Page 31 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

THRUST LEVER Identification (1.A.a) The thrust lever assembly is located in the cockpit. The thrust lever provides a dual RVDT interface to the FADEC for the thrust lever position. The FADEC provides excitation and demodulation of the RVDT’s. The thrust levers move over a sector divided into five areas separated by unique positions. The five are Max Reverse, Reverse Idle, Idle, TO/GA and Max. The Thrust Lever incorporates the following design details: - Single thrust lever per engine for both reverse and forward thrust control - Thrust lever trigger is to prevent reverse thrust range selection in flight - One mechanical detent for TOGA thrust - Auto Throttle system disconnect switch Refer to figure, THRUST LEVER. Purpose (2.B.b) The thrust lever assembly helps in scheduling of forward and reverse thrust for each engine. 2 3 0 1 5 0 0 0 0 0 3 7 D S

EFFECTIVITY ALL

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Page 32 Aug 05

GE AIRCRAFT ENGINES

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CF34-10E

TRAINING MANUAL

THRUST LEVER QUADRANT ANGLES AND DETENTS

EFFECTIVITY ALL

73-00-00


FUEL AND CONTROL

Page 33 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

AUTO THROTTLE INTERFACE Identification (1.A.a) The thrust lever position is controlled by position control loop closed around the thrust lever via an Auto Throttle (AT) system and lever position feedback sensor (RVDT) output. Purpose (2.B.b) A/T interface is responsible for maintaining relationship between engine thrust and the thrust lever position.

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Operation (3.C.c) The Thrust Control Quadrant (TCQ), electronics assist the A/T system by providing necessary rate and torque control for each of the servo drives. The TCQ control electronics provide full capability of receiving and transmitting data for the ARINC 429 system format. Rate feedback of each thrust lever and engage (disconnect) status of TCQ is also transmitted back to the A/T system via the ARINC 429 buss. The TCQ control electronics receive, from the A/T system, a rate command for each thrust lever servomotor. Rate commands are used by TCQ control electronics to provide closed rate loop around each thrust lever servomotor. Hall effect sensors located in servomotors provide necessary rate feedback information. The TCQ control electronics also provide internal Pulse Width Modulated (PWM) motor current loop, to regulate the torque produced by the servomotors.

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 34 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

5 3 0 1 1 0 0 0 0 0 3 7 D S

EFFECTIVITY ALL

73-00-00


FUEL AND CONTROL

Page 35 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE FUEL CONTROL SYSTEM Identification (1.A.a) The CF34-10E engine control system is a computer based electronic engine control system. It is composed of a two channel Full Authority Digital Engine Control (FADEC), a Fuel Metering Unit (FMU), a Permanent Magnet Alternator (PMA), engine sensors, Variable Stator Vane (VSV) actuator, Variable Bleed Valve (VBV) actuator, Transient Bleed Valve (TBV) actuator, High Pressure Turbine Clearance Control Valve (HPTCC) Actuator, Low Pressure Turbine Clearance Control Valve (LPTCC) and an ignition system for each engine. The system controls the engine in response to thrust command inputs from the aircraft and provides information to the aircraft for cockpit indication, maintenance reporting and engine condition monitoring.

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Purpose (2.B.b) The control system provides steady state and transient regulation of thrust while protecting the following engine limits: compressor aerodynamic stability limits, compressor aeromechanical limits, combustor blowout limits, and fan and core rotor mechanical limits. The control also provides independent protection of core rotor over speed limits. The control does not provide explicit protection against a turbine temperature exceedance with the exception of ground starts.

FADEC channel operates as the “standby” FADEC channel processing all inputs and software, however the electronic control outputs (except over speed solenoid driver) are disabled during normal engine operation. In addition, the “standby” FADEC channel shares selected sensor inputs, airframe commands, and FADEC status information using a cross channel serial data bus in order to maintain the maximum system fault tolerance. During operation with two capable FADEC channels, in control software logic will cause the FADEC channels to alternate control on each successive engine start. The FADEC power supply is primarily provided by the PMA during engine operation with a 28 VDC airframe input for starting and backup.

Operation (3.C.c) One FADEC channel operates as the “in control” FADEC channel providing electronic control outputs. The other

EFFECTIVITY ALL

73-00-00


FUEL AND CONTROL

Page 36 May 07

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

F M U DP3

ENGINE FUEL CONTROL SYSTEM

EFFECTIVITY ALL

73-00-00


FUEL AND CONTROL

Page 37 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

9 3 0 1 1 0 1 0 1 2 3 7 D S

EFFECTIVITY ALL

73-21-01


FUEL AND CONTROL

Page 39 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

engine permanent magnet alternator (PMA) or by aircraft 28 VDC power. During normal operation at and above engine idle speeds, the FADEC will use PMA power. During starting and sub idle operation, the FADEC will use aircraft 28 VDC power.

0 4 0 1 5 0 1 0 1 2 3 7 D S

EFFECTIVITY ALL

73-21-01


FUEL AND CONTROL

Page 40 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL


EFFECTIVITY ALL

73-21-01


FUEL AND CONTROL

Page 41 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

T12 SENSOR Identification (1.A.a) The engine T12 sensor is mounted in the flow stream in front of the fan and well above the engine centerline with one element hardwired to each FADEC channel. The T12 sensor is a dual element Resistive Thermal Device (RTD). Refer to figure, T12 SENSOR. Purpose (2.B.b) The T12 sensor provides the total ambient temperature to the FADEC for its various calculations. Note: Each FADEC channel uses four sources of engine inlet temperature data: two engine sensors (T12) and two aircraft Air Data System signals (TAT).

2 4 0 1 5 0 1 0 1 2 7 7 D S

EFFECTIVITY ALL

77-21-01


FUEL AND CONTROL

Page 42 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

3 4 0 1 1 0 1 0 1 2 7 7 D S

EFFECTIVITY ALL

77-21-01


FUEL AND CONTROL

Page 43 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

T25 SENSOR Identification (1.A.a) The engine T25 sensor is mounted in the flow stream in front of the compressor with one element hardwired to each FADEC channel. The T25 sensor is a dual element Resistive Thermal Device (RTD). Refer to figure, T25 SENSOR. Purpose (2.B.b) The T25 sensor measures and sends the compressor inlet air temperature to the FADEC.

4 4 0 1 5 0 4 0 1 2 7 7 D S

EFFECTIVITY ALL

77-21-04


FUEL AND CONTROL

Page 44 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

5 4 0 1 1 0 4 0 1 2 7 7 D S

EFFECTIVITY ALL

77-21-04


FUEL AND CONTROL

Page 45 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

T3 SENSOR Identification (1.A.a) The T3 sensor is mounted on the combustor frame. The T3 sensor is a dual element thermocouple. On engine serial numbers 994577 and above this sensor was removed from production. The FADEC replaces the sensor using algorithms. Refer to figure, T3 SENSOR. Purpose (2.B.b) The T3 sensor measures the compressor discharge air temperature and provides it to FADEC.

6 4 0 1 5 0 5 0 1 2 7 7 D S

EFFECTIVITY ALL

77-21-05


FUEL AND CONTROL

Page 46 Feb 09

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

• The T3 sensor has been removed as an input and replaced by algorithms within the FADEC. • SB 73-0017 – Software load 5.32, sensor input change. • SB 72-113 and 114 – Introduction of new harnesses. • SB 72-115 – Removal of the sensor from the engine.

7 4 0 1 1 0 5 0 1 2 7 7 D S

EFFECTIVITY ALL

77-21-05


FUEL AND CONTROL

Page 47 Feb 09

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

DELTA P3 SENSOR Identification (1.A.a) The delta P3 sensor is at approximately 6:30 position aft of the fan case. The delta P3 sensor is a pressure transducer. Refer to figure, DELTA P3 SENSOR. Purpose (2.B.b) The delta P3 sensor measures the High Pressure Compressor (HPC) ninth stage bleed pressure and sends the signal to the Full Authority Digital Engine Control (FADEC). This information is used by the FADEC for making fuel schedule adjustments. This helps in improving engine acceleration.

8 4 0 1 5 0 5 0 1 2 7 7 D S

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 48 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Ninth Stage Bleed Fitting to ECS System

PS3 Tube: DP Low Side

PS3 Tube: DP High Side LO HI

PS3 Tube: DP Low Side

PS3 Tube: DP High Side

9 4 0 1 1 0 5 0 1 2 7 7 D S

Channels “A&B” To FADEC

DELTA P3 SENSOR EFFECTIVITY ALL

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FUEL AND CONTROL

Page 49 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

TCASE SENSOR Identification (1.A.a) The Tcase sensor is on the High Pressure Turbine (HPT) manifold. The Tcase sensor consists of a single type K thermocouple (TC). On engine serial numbers 994577 and above this sensor was removed from production. The FADEC replaces the sensor using algorithms. Refer to figure, TCASE SENSOR. Purpose (2.B.b) The Tcase sensor measures HPT case shroud temperature and is an input to the HPT clearance control algorithm.

0 5 0 1 5 0 5 0 1 2 7 7 D S

EFFECTIVITY ALL

77-21-05


FUEL AND CONTROL

Page 50 Feb 09

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

• The TCase sensor has been removed as an input and replaced by algorithms within the FADEC. • SB 73-0017 – Software load 5.32, sensor input change. • SB 72-113 and 114 – Introduction of new harnesses. • SB 72-115 – Removal of the sensor from the engine.

1 5 0 1 1 0 6 0 1 2 7 7 D S

EFFECTIVITY ALL

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FUEL AND CONTROL

Page 51 Feb 09

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FAN SPEED SENSOR - N1 Identification (1.A.a) The fan speed sensor (N1) is on the fan case at the 3 o’clock position. The N1 sensor is a three coil reluctance transmitter assembly that provides signals indicative of an aircraft engine fan rotational speed. The N1 sensor gets its signal from a 25 tooth wheel on the fan shaft. Refer to figure, FAN SPEED SENSOR - N1. Purpose (2.B.b) The N1 sensor provides two fan rotor speed measurements to the FADEC and one output that is used in the Engine Vibration Monitoring System.

2 5 0 1 5 0 1 0 1 1 7 7 D S

EFFECTIVITY ALL

77-11-01


FUEL AND CONTROL

Page 52 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

3 5 0 1 1 0 1 0 1 1 7 7 D S

EFFECTIVITY ALL

77-11-01


FUEL AND CONTROL

Page 53 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

CORE SPEED SENSOR - N2 Identification (1.A.a) The core speed (N2) sensor is on the Accessory Gearbox (AGB). The N2 sensor is a two coil reluctance transmitter assembly. Refer to figure, CORE SPEED SENSOR - N2. Purpose (2.B.b) The N2 sensor provides the core speed signal to each channel of the FADEC.

4 5 0 1 5 0 2 0 1 1 7 7 D S

EFFECTIVITY ALL

77-11-02


FUEL AND CONTROL

Page 54 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

5 5 0 1 1 0 2 0 1 1 7 7 D S

EFFECTIVITY ALL

77-11-02


FUEL AND CONTROL

Page 55 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

PERMANENT MAGNET ALTERNATOR (PMA) Identification (1.A.a) The Permanent Magnetic Alternator (PMA) is mounted on the engine AGB. Refer to figure, PERMANENT MAGNET ALTERNATOR PMA. Purpose (2.B.b) The purpose of the PMA is to provide power to the FADEC. The PMA also provides N2 signals for the engine control system as well as an N2 signal for use in the Engine Vibration Monitoring System. (EVM) The signal is generated by the rotational speed of the PMA. Operation (3.C.c) The selection between the aircraft 28 VDC power supply and the engine supplied PMA power supply is performed automatically by the FADEC. Each FADEC channel has a dedicated input from the engine PMA. When the engine speed is greater than 50% N2, the dedicated alternator input has the capability to provide all electrical power for the FADEC system. 6 5 0 1 5 0 0 0 1 1 7 7 D S

EFFECTIVITY ALL

77-11-00


FUEL AND CONTROL

Page 56 May 07

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

and EVM 7 5 0 1 1 0 0 0 1 1 7 7 D S

EFFECTIVITY ALL

77-11-00


FUEL AND CONTROL

Page 57 May 07

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE CONFIGURATION PLUG Identification (1.A.a) The engine configuration plug (ECP) is on the fan case, mounted to the FMU bracket. Refer to figure, ENGINE CONFIGURATION PLUG. Purpose (2.B.b) The ECP provides data that is read by the FADEC at the time of power up on the ground. The ECP provides information on the engine serial number, engine thrust rating, hardware configuration, and N1 modification (N1 trim). Operation (3.C.c) The FADEC will validate the engine configuration by comparing the engine configuration from the ECP to the FADEC software.

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Similarly, the FADEC will validate the aircraft application by comparing the aircraft application from the Application Identification (APPID) to the FADEC software. If the FADEC Cannot accommodate the aircraft application, the FADEC will transmit the appropriate EICAS messages to the cockpit. The aircraft MAU will also compare the engine ratings from both engines as a validity check. Each engine will provide an ECP engine rating from each FADEC channel.

EFFECTIVITY ALL

73-21-07


FUEL AND CONTROL

Page 58 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

9 5 0 1 1 0 7 0 1 2 3 7 D S

EFFECTIVITY ALL

73-21-07


FUEL AND CONTROL

Page 59 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

IGNITION SYSTEM


74-00-00 IGNITION SYSTEM

Page 1 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

IGNITION SYSTEM Objectives Given an objective exercise, the student will identify: - Selected components of the ignition system (1.A.a) - The ignition system distribution (1.A.a) - The ignition power of the ignition system (1.A.a) - The location of the ignition exciter (1.A.a) - The location of the ignition leads (1.A.a) - The location of the spark igniters (1.A.a) - The ignition system control (1.A.a) Given an objective exercise, the student will select the purpose of: - Selected components of the ignition system (2.B.b) - The ignition system distribution (2.B.b) - The ignition power of the ignition system (2.B.b) - The ignition exciter (2.B.b) - The ignition leads (2.B.b) - The spark igniters (2.B.b) - The ignition system control (2.B.b)

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Given an objective exercise, the student will select the operation of: - The ignition exciter (3.C.c)



Page 2 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL




Page 3 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

IGNITION SYSTEM COMPONENTS Identification (1.A.a) Each engine has two ignition systems that operate separately. The ignition system operates manually or automatically . The engine ignition system has the parts that follow: - Ignition exciters - Ignition leads - Spark igniters Purpose (2.B.b) The engine ignition system supplies the necessary spark to burn the air/fuel mixture during normal ground starts, and ignition ON operating conditions such as landing and takeoff in bad weather.

4 0 0 1 5 0 0 0 0 0 4 7 D S



Page 4 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL




Page 5 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

IGNITION SYSTEM DISTRIBUTION Identification (1.A.a) The ignition system supplies high-voltage electrical energy for engine start in normal conditions or abnormal engine conditions to prevent engine flameout. Purpose (2.B.b) The Modular Avionics Units (MAU) interface with a three position switch (OFF/AUTO/OVRD) that allows the pilots to send ignition commands to the FADEC. The FADEC receives the ignition switch position from the MAUs and determines the required ignition command. The FADEC sends the potential to the two ignition exciters that rectify, step up and send voltage to the spark igniters.

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The switch positions will have the following actions: OFF – FADEC commands commands ignition ignition “OFF” AUTO – FADEC will enable enable either either A or B ignition “ON” “ON” during ground starts. Ignition A&B system will alternate on each ground start as commanded by the FADEC. OVRD – Commands ignition A and B “ON” through FADEC and and by directly powering the relay coils located in the Emergency Integrated Control Center (EICC) based on condition. Which igniter igniter is “ON” is PS3.0 dependent dependent:: • Below 170 psia PS3.0 PS3.0 – Both A and B will will be commanded commanded “ON” “ON” through FADEC and override path. • Above 170 170 psia PS3.0 PS3.0 – Ignition A only will be commande commanded d “ON” independent independent of FADEC. This is done to help prevent damage to ignition leads due to the high pressure condition.



Page 6 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

IGNITION SWITCH PANEL



Page 7 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

IGNITION POWER Identification (1.A.a) Alternating current (AC) power (115 volt) supplied by the aircraft electrical system is sent to both ignition exciters. Exciter A receives 115 VAC 400 HZ from the aircraft via the Standby AC Bus. Exciter B receives power from the aircraft AC1 (engine 1) and AC2 (engine 2) Bus. Purpose (2.B.b) The exciters use this supply voltage to the ignition exciters where it is converted to direct current (DC), capacitance discharge. Ignition system selection is displayed on EICAS as follows: IGN A – FADEC has commanded commanded igniter A to be energized energized due to a ground start. Green indication. IGN B – FADEC has commanded commanded igniter B to be energized energized due to a ground start. Green indication. 8 0 0 1 5 0 0 0 0 0 4 7 D S

IGN A B – FADEC has commanded commanded igniters A and B to be energized due to an in flight start or an auto relight. Green indication. IGN OFF – FADEC has locked ignition ignition off due to pilot procedure procedure to dry motor or fire handle has been activated. Cyan indication.



Page 8 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

IGN A or B - Green Green - Normal Normal groun ground d start start IGN A B - Green Green – Auto-rel Auto-religh ightt or in flight flight start start IGN OFF - Cyan Cyan – Ignitio Ignition n locked locked OFF for dry motor or fire handle pulled

IGNITION POWER INDICATION



Page 9 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

IGNITION EXITER Identification (1.A.a) There are two interchangeable exciters mounted at 6 o'clock on the fan case. Each exciter serves one igniter. Purpose (2.B.b) The exciters convert aircraft 115 AC voltage to DC voltage, providing 14,000 to 18,000 VDC capacitance discharge at a rate of one pulse per second to the igniter plugs.

Operation (3.C.c) The ignition exciter units change, rectify, and store electrical energy in a capacitor. The capacitor then sends an electrical pulse to the spark igniters. For safety, a bleed down resister is provided to dissipate any residual charge from the capacitor.

0 1 0 0 1 5 0 0 0 0 0 4 7 D S



Page 10 Aug 05

GE AIRCRAFT ENGINES

1 1 0 0 1 1 0 0 0 0 0 4 7 D S

CF34-10E

TRAINING MANUAL

IGNITION EXCITER



Page 11 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

IGNITION LEADS Identification (1.A.a) The ignition leads connects between the ignition exciters and the igniters. Fan air, supplied from upstream of the LPTACC valve, is used to cool the ignition lead from the six o’clock mast position to the igniter. Purpose (2.B.b) The ignition leads carry the high voltage electrical power from the ignition exciters to the spark igniters. Maintenance Note: When removing the ignition lead, do not twist the lead. Over time, the diolectric (contact) becomes brittle and can be damaged.

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Page 12 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

(Diolectric)

Cooling Airflow 3 1 0 0 1 1 0 0 0 0 0 4 7 D S



Page 13 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

SPARK IGNITERS Identification (1.A.a) There are two spark igniters. The left igniter is on the combustion chamber at 8 o'clock. The right igniter is on the combustion chamber at 4 o'clock. The igniter consists of a center and outer electrode. A semiconductor surface coats the tip between the two electrodes. The tip extends inside the combustion liner, exposed to the fuel/air mixture. Purpose (2.B.b) Each igniter provides the electrical spark needed to start or maintain combustion.

4 1 0 0 1 5 0 0 0 0 0 4 7 D S



Page 14 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

5 1 0 0 1 1 0 0 0 0 0 4 7 D S



Page 15 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE 1 AND 2 EXCITER A CMD

.

..

OVRD

OVRD

.

7 1 0 1 1 0 0 0 0 0 4 7 D S

ENGINE 1 AND 2 IGNITION OVERRIDE

IGNITION SYSTEM CONTROL EFFECTIVITY ALL GE PROPRIETARY INFORMATION


Page 17 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE AIR SYSTEM

EFFECTIVITY ALL

75-00-00


ENGINE AIR SYSTEM

Page 1 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE AIR SYSTEM Objectives Given an objective exercise, the student will identify: - Selected engine airflow systems (1.A.a) - The compressor airflow control (1.A.a) - The variable bleed valve (VBV) subsystem (1.A.a) - The variable stator valve (VSV) subsystem (1.A.a) - The transient bleed valve (TBV) subsystem (1.A.a) - The HPT active clearance control (HPTACC) subsystem (1.A.a) - The LPT active clearance control (LPTACC) subsystem (1.A.a) Given an objective exercise, the student will select the purpose of: - Selected engine airflow systems (2.B.b) - The compressor airflow control (2.B.b) - The variable bleed valve (VBV) subsystem (2.B.b) - The variable stator valve (VSV) subsystem (2.B.b) - The transient bleed valve (TBV) subsystem (2.B.b) - The HPT active clearance control (HPTACC) subsystem (2.B.b) - The LPT active clearance control (LPTACC) subsystem (2.B.b)

2 0 0 1 5 0 0 0 0 0 5 7 D S

Given an objective exercise, the student will select the operation of: - The compressor airflow control (3.C.c) - The variable bleed valve (VBV) subsystem (3.C.c) - The variable stator valve (VSV) subsystem (3.C.c) - The transient bleed valve (TBV) subsystem (3.C.c) - The HPT active clearance control (HPTACC) subsystem (3.C.c) - The LPT active clearance control (LPTACC) subsystem (3.C.c)

EFFECTIVITY ALL

75-00-00


ENGINE AIR SYSTEM

Page 2 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL


EFFECTIVITY ALL

75-00-00


ENGINE AIR SYSTEM

Page 3 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE AIRFLOW SYSTEMS Identification (1.A.a) The engine airflow systems consists of the following subsystems: - Compressor airflow control - Variable bleed valve (VBV) subsystem - Variable stator vanes (VSV) subsystem - Transient bleed valve (TBV) subsystem - High pressure turbine active clearance control (HPTACC) subsystem - Low pressure turbine active clearance control (LPTACC) subsystem Purpose (2.B.b) The engine airflow system controls the amount of air passing through the compressor. It also controls the clearance between the turbine blade and the turbine shrouds, for both the HPT and the LPT.

4 0 0 1 5 0 0 0 0 0 5 7 D S

EFFECTIVITY ALL

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ENGINE AIR SYSTEM

Page 4 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

LPTACC

VBV

HPTACC

VSV TBV ENGINE AIRFLOW SYSTEMS EFFECTIVITY ALL

75-00-00


ENGINE AIR SYSTEM

Page 5 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

VARIABLE BLEED VALVE (VBV) SUBSYSTEM Identification (1.A.a) The variable bleed valve (VBV) subsystem is comprised of a dual-coil, two-stage electrohydraulic servo valve (EHSV) which is integral to the FMU, two fuel-driven actuators, and the FADEC. Each actuator includes a single-coil linear variable differential transducer (LVDT) that provides actuator position feedback to each FADEC channel. Refer to figure, VARIABLE BLEED SYSTEM. Purpose (2.B.b) The VBV subsystem helps in controlling the booster operating line to provide optimum booster performance at steady state and prevent stalls.

linkages on the doors to a variable open or closed position. The VBV actuator position demand is computed in the FADEC software to optimize the position of the variable bleed valves as a function of the current steady-state and transient engine operating condition. Primarily, the VBV position demand is computed as a function of corrected N1. The position demand is modified during transient operation to maintain booster operability margins.

Operation (3.C.c) During normal operation, the VBV actuators are positioned by the FADEC in a closed-loop fashion via a current command to the VBV EHSV. The LVDT feedback is used to close the position loop through the FADEC. 6 0 0 1 5 0 0 0 0 0 5 7 D S

The position is controlled by the FADEC as a function of N1 and N2. The VBV doors operate from idle, fully open, to take off power, fully closed. During reverse thrust, the FADEC will partially open the doors to increase booster stability and help prevent any core engine foreign object damage. (FOD) Two VBV actuators position bellcranks attached to a 360 degree unison ring that moves

EFFECTIVITY ALL

75-00-00


ENGINE AIR SYSTEM

Page 6 Sept 05

GE AIRCRAFT ENGINES

CF34-10E

Open

TRAINING MANUAL

N1K

VBV Channel A Channel B

Closed

FADEC Position Demand

N2K FMU

LVDT Feedback

Bellcrank and Unison Ring

LVDT Feedback Fuel Pressure Lines From FMU (Rod/Head)

7 0 0 1 1 0 0 0 0 0 5 7 D S

Variable Bleed System (VBV) EFFECTIVITY ALL

75-00-00


ENGINE AIR SYSTEM

Page 7 May 07

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

VARIABLE STATOR VANES (VSV) SUBSYSTEM Identification (1.A.a) The variable stator vane (VSV) subsystem is comprised of a dual-coil, two-stage EHSV which is integral to the FMU, two fuel-driven actuators, and the FADEC. Each actuator includes a single-coil LVDT that provides actuator position feedback to each FADEC channel. Refer to figure, VARIABLE STATOR VANE (VSV) SYSTEM.

compressor stators as a function of the current steady-state and transient engine operating condition. Primarily, the VSV position demand is computed as a function of corrected N2. The position demand is modified during transient operation and during combustor relights to overclose the stators and maintain compressor operability margins.

Purpose (2.B.b) The VSV system controls the amount of air passing through the high pressure compressor. This helps in better fuel efficiency at all power settings as well as preventing compressor surges and stalls. Operation (3.C.c) During normal operation, the VSV actuators are positioned by the FADEC in a closed-loop fashion via a current command to the VSV EHSV. The LVDT feedback is used to close the position loop through the FADEC. 8 0 0 1 5 0 0 0 0 0 5 7 D S

The VSV system varies the angle-of-attack of the variable HPC stator vanes through two VSV actuators attached to torsion tubes and unison rings which rotate all stages together. The VSV actuator position demand is computed in the FADEC software to optimize the position of the

EFFECTIVITY ALL

75-00-00


ENGINE AIR SYSTEM

Page 8 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Closed FADEC

VSV Channel A

Steady State

Open

Channel B

N2K Position Demand

FMU Fuel Pressure Lines From FMU (Rod/Head) LVDT Feedback 9 0 0 1 1 0 0 0 0 0 5 7 D S

Variable Stator Vane System (VSV)

EFFECTIVITY ALL

75-00-00


ENGINE AIR SYSTEM

Page 9 May 07

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

TRANSIENT BLEED VALVE (TBV) SUBSYSTEM Identification (1.A.a) The transient bleed valve (TBV) subsystem is comprised of the TBV, the fuel metering unit (FMU), and the FADEC. The FMU includes a dual-coil, two-stage electrohydraulic servo valve (EHSV), which is used to position the TBV. The TBV has an actuator which positions a butterfly valve and a dual-coil LVDT that provides actuator position feedback to the FADEC.

and engine starting, the TBV is scheduled open in order to bleed the compressor and maintain compressor operability margins by bleeding compressor discharge pressure (CDP) air into the low pressure turbine (LPT) cooling circuit.

Refer to figure, TRANSIENT BLEED VALVE (TBV) SYSTEM. Purpose (2.B.b) The TBV subsystem helps maintain compressor operability margins. Operation (3.C.c) During normal operation, the TBV is positioned by the FADEC in a closed-loop fashion via a current command to the TBV electro EHSV in the FMU. The LVDT feedback is used to close the position loop through the FADEC. 0 1 0 1 5 0 0 0 0 0 5 7 D S

The TBV position loop demand is computed in the FADEC software to optimize the compressor discharge bleed from the engine as a function of the current transient engine operating condition. During steady-state engine operation, the TBV will be commanded closed. During transient operation, combustor relights,

EFFECTIVITY ALL

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ENGINE AIR SYSTEM

Page 10 Sept 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FADEC

FMU

Position Demand

Valve Discharge into LPT Nozzle Fuel Pressure Lines From FMU

Ninth Stage Air

Transient Bleed Valve 1 1 0 1 1 0 0 0 0 0 5 7 D S

LVDT Feedback

Channel A Channel B

Transient Bleed Valve System (TBV) EFFECTIVITY ALL

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ENGINE AIR SYSTEM

Page 11 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

HPT ACTIVE CLEARANCE CONTROL (HPTACC) SUB-SYSTEM

Identification (1.A.a) The high pressure turbine active clearance control (HPTACC) subsystem is comprised of the HPTACC valve, the FMU, and the FADEC. The FMU includes a dual-coil, two-stage electro hydraulic servo valve (EHSV) which positions the HPTACC valve actuator. The HPTACC has an actuator which positions dual butterfly valves and a dual-coil LVDT that provides actuator position feedback to the FADEC.

through individual valve elements positioned by a signal fuel actuator. The air mixes downstream in the air duct and is directed to the turbine clearance control cavity, which distributes the air to the shroud segments. The air then exits into the LPT stage one nozzle cooling circuit. The position demand is computed in the FADEC software to optimize the turbine clearance as a function of the engine operating conditions.

Refer to figure, HPT ACTIVE CLEARANCE CONTROL (HPTACC). Purpose (2.B.b) The HPTACC subsystem helps in minimizing the clearance between the HPT blades and the HPT shrouds.

2 1 0 1 5 0 0 0 0 0 5 7 D S

Operation (3.C.c) During normal operation, the HPTACC valve is positioned by the FADEC in a closed-loop fashion via a current command to the HPTACC EHSV in the FMU. The LVDT feedback is used to close the position loop through the FADEC. The HPTACC valve is positioned by the FADEC based on representative FADEC algorithms of Tcase and T3. Both fourth and ninth stage air enters the valve where it is metered

EFFECTIVITY ALL

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ENGINE AIR SYSTEM

Page 12 Feb 089

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FADEC

Mixed Supply to HPT Case Fuel Pressure Lines From FMU

FMU

9TH Stage Supply

Position Demand

HPTCC VALVE

Channel A Channel B LVDT Feedback

3 1 0 1 1 0 0 0 0 0 5 7 D S

4TH Stage Supply

HPT Active Clearance Control System (HPTACC) EFFECTIVITY ALL

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ENGINE AIR SYSTEM

Page 13 Feb 09

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

LPT ACTIVE CLEARANCE CONTROL (LPTACC) SUB-SYSTEM

Identification (1.A.a) The low pressure turbine active clearance control (LPTACC) subsystem is comprised of the LPTACC valve and the FADEC. The LPTACC valve is self-modulating which positions a butterfly valve and a dual-coil LVDT that provides actuator position feedback to the FADEC. The valve is located at the 5:00 o’clock position on the engine core. Refer to figure, LPT ACTIVE CLEARANCE CONTROL (LPTACC). Purpose (2.B.b) The LPTACC subsystem helps in minimizing the clearance between the LPT blades and the LPT shrouds. Operation (3.C.c) During normal operation, the LPTACC valve is positioned by the FADEC based on ITT and fan speed. The LVDT feedback is used to close the position loop through the FADEC. The position demand is computed in the FADEC software to optimize the turbine clearance as a function of the engine operating conditions. Cooling air is supplied from a fan discharge scoop in the extension ring at the 6:00 o’clock position, it then passes through the control valve and impinges on the LPT through a manifold that surrounds the case. The valve will be at a minimum flow condition during idle and takeoff to allow stator case growth and eliminate hard rubs in the honeycomb seals. The valve will be at high flow during cruise to minimize case growth and maintain clearances.

EFFECTIVITY ALL

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ENGINE AIR SYSTEM

Page 14 March 06

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

Fuel Pressure LPT Cooling Manifold FADEC

Position Demand

Channel A, B Feedback

LPTACC Control Valve

Fan Air Inlet

Igniter Cooling Air Supply

LPT ACTIVE CLEARANCE CONTROL (LPTACC) EFFECTIVITY ALL

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ENGINE AIR SYSTEM

Page 15 May 07

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE INDICATING

EFFECTIVITY ALL

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ENGINE INDICATING

Page 1 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE INDICATING Objectives Given an objective exercise, the student will identify: - The indicating system of the CF34-10E engine (1.A.a) - Fan speed indication – N1 (1.A.a) - N1 exceedance indication (1.A.a) - Core speed indication – N2 (1.A.a) - N2 exceedance indication (1.A.a) - Inter Turbine Temperature indication – ITT (1.A.a) - ITT exceedance indication (1.A.a) - Vibration indication system (1.A.a) - Accelerometer (1.A.a) - AVM signal conditioner (1.A.a) - Oil filter bypass/low oil PS indication (1.A.a) - Oil Temperature/Quantity (1.A.a) - Fuel filter bypass/low fuel PS indication (1.A.a)

2 0 0 1 5 0 0 0 0 0 7 7 D S

Given an objective exercise, the student will select the purpose of: - The indicating system of the CF34-10E engine (2.B.b) - Fan speed indication – N1 (2.B.b) - N1 exceedance indication (2.B.b) - Core speed indication – N2 (2.B.b) - N2 exceedance indication (2.B.b) - Exhaust gas temperature indication – EGT (2.B.b) - ITT exceedance indication (2.B.b) - Vibration indication system (2.B.b) - Accelerometer (2.B.b) - AVM signal conditioner (2.B.b) - Oil filter bypass/low oil PS indication (2.B.b)

EFFECTIVITY ALL

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ENGINE INDICATING

Page 2 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

- Oil Temperature/Quantity (2.B.b) - Fuel filter bypass/low fuel PS indication (2.B.b) Given an objective exercise, the student will select the operation of: - Fan speed indication – N1 (3.C.c) - Core speed indication – N2 (3.C.c) - Inter Turbine Temperature indication – ITT (3.C.c) - Vibration indication system (3.C.c) - Accelerometer (3.C.c)

3 0 0 1 5 0 0 0 0 0 7 7 D S

EFFECTIVITY ALL

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ENGINE INDICATING

Page 3 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

INDICATING SYSTEM Identification (1.A.a) The engine indicating system includes the components that follows: - Engine Indication Cockpit Annunciation System (EICAS) - Full Authority Digital Engine Control (FADEC) - Modular Avionics Unit (MAU) - Multi Function Display (MFD) - Central Maintenance Computer (CMC) The engine indicating system includes the subsystem components that follow: - Fan speed indication (N1) - Core speed indication (N2) - Inter Turbine Temperature (ITT) - Vibration indicating - Oil filter bypass/low oil PS - Oil Temperature/Quantity - Fuel filter bypass/low fuel PS

4 0 0 1 5 0 0 0 1 1 7 7 D S

The MAU will compare engine parameters with their limits and indicate an exceedence to the cockpit by changing the color indication. If an amber line is exceeded, the indication is to turn amber, if a red line is exceeded, the indication is to turn red. The MAU will transmit oil temperature, oil level, oil pressure and N1 and N2 vibration to the FADEC to allow limit tracking. These parameters are initially sent through the MAU and not FADEC. The FADEC will then test these signals against the limits. The MFD will allow for resetting of the exceedence to the FADEC. The CMC is used for storage and transfer of exceedence data.

Purpose (2.B.b) The engine indicating system operates with the engine systems to supply data displays to EICAS. Any exceedance values are set by the FADEC and will make a determination of Amberline and Redline engine limits. The value of the limits may vary based on engine operating condition and are transmitted to the aircraft for use in setting cockpit EICAS displays.

EFFECTIVITY ALL

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ENGINE INDICATING

Page 4 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

5 0 0 1 1 0 0 0 1 1 7 7 D S

EFFECTIVITY ALL

77-11-00


ENGINE INDICATING

Page 5 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FAN SPEED INDICATION - N1 Identification (1.A.a) The fan speed (N1) indicating system consists of fan speed sensor. The N1 sensor is on the fan case at the 3 o’clock position. Purpose (2.B.b) The N1 sensor sends the fan speed signal to these components: - Full authority digital engine control (FADEC) - Engine Indicating Cockpit Annunciation System (EICAS) - Engine vibration monitoring (EVM)

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Operation (3.C.c) The N1 speed sensor sends an analog signal to the FADEC. The FADEC changes these signals to digital and sends the signals (channel A and channel B) to the EICAS. The EICAS uses the signals from the FADEC to show N1 on the EICAS display system and range from 0% to a maximum displayed value of 110%. Selection Logic: If both the local and cross talk N1 signals are invalid in the channel in control, then a modeled N1 value is selected and the engine will run on a N2 shadow governor resulting in a thrust change up to +/- 10%. Also, the N1 signal transmitted to the other engine FADEC will be set to zero. The engine vibration monitoring system (EVM) uses input from the N1 speed sensor to help calculate vibration levels.

EFFECTIVITY ALL

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ENGINE INDICATING

Page 6 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

N1 Signal N1 Vib Level

7 0 0 1 1 0 0 0 0 0 7 7 D S

N1 SENSOR

N1 INDICATING EFFECTIVITY ALL

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ENGINE INDICATING

Page 7 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

N1 EXCEEDANCE INDICATION Identification (1.A.a) The N1 EXCEEDANCE is indicated by a change of color on the N1 dial and digital readout display of the EICAS. Purpose (2.B.b) This indicates that N1 exceedance has occurred. As exceedence speed increases, the indication will go from an amber warning to a red indication.

8 0 0 1 5 0 0 0 0 0 7 7 D S

EFFECTIVITY ALL

77-00-00


ENGINE INDICATING

Page 8 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

EXCEEDANCE

NORMAL

N1 Signal

EICAS Cyan bug to indicate N1 Rating following the engine thrust rating annunciation. EXCEEDANCE THRESHOLD INDICATED BY A RED TICK MARK

9 0 0 1 1 0 0 0 0 0 7 7 D S

N1 SENSOR

N1 EXCEEDANCE EFFECTIVITY ALL

77-00-00


ENGINE INDICATING

Page 9 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

CORE SPEED INDICATION - N2 Identification (1.A.a) The core speed (N2) indicating system consists of a core speed sensor. The core speed (N2) sensor is on the accessory gearbox (AGB).There are also two N2 signals, one per channel, sent from the permanent magnet alternator (PMA) for use in monitoring N2 speed. Purpose (2.B.b) The N2 sensor sends the core speed signal to these components: - Full authority digital engine control (FADEC) - EICAS - MAU - Engine vibration monitoring (EVM)

0 1 0 1 5 0 0 0 0 0 0 7 D S

Operation (3.C.c) Each FADEC channel receives an N2 signal from a magnetic reluctance pickup that reads 47 teeth on a special gear on the hydraulic pump shaft. Each channel also determines N2 from the three phase alternator windings. Each hardwired N2 signal is shared through the CCDL so each channel receives four independent electrical core speed inputs. The FADEC changes these signals to digital and sends the signals (channel A and channel B) to the MAUs. The MAUs use the signals from the FADEC to show N2 on the EICAS. The engine vibration monitoring system (EVM) uses analog input from the speed sensor to help calculate vibration levels.

EFFECTIVITY ALL

77-00-00


ENGINE INDICATING

Page 10 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

1 1 0 1 1 0 0 0 0 0 0 7 D S

EFFECTIVITY ALL

77-00-00


ENGINE INDICATING

Page 11 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

N2 EXCEEDANCE INDICATION Identification (1.A.a) The N2 EXCEEDANCE is indicated by a change of color on the N2 dial and digital readout display of the EICAS. Purpose (2.B.b) This indicates that N2 exceedance has occurred. As exceedence speed increases, the indication will go from white status to a red indication. Core speeds will read out status up to 100% (18018 N2). Overspeed trip level is set at 100.95% (18190 N2)

2 1 0 1 5 0 0 0 0 0 0 7 D S

EFFECTIVITY ALL

77-00-00


ENGINE INDICATING

Page 12 Mar 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

N2 EXCEEDENCE RED INDICATION AT 18018 RPM (100%) OVERSPEED TRIP AT 18190 RPM (100.95%)

EXCEEDENCE

EFFECTIVITY ALL

77-00-00


ENGINE INDICATING

Page 13 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

INTER TURBINE TEMPERATURE INDICATION - ITT Identification (1.A.a) The inter turbine temperature (ITT) indicating system includes the components that follow: - Nine ITT probes - Three ITT thermocouple lead assemblies The ITT probes are inside the second stage nozzles of the low pressure turbine (LPT). A wire harness connects the full authority digital engine control (FADEC) to the junction box near the ITT probes. Purpose (2.B.b) The ITT indicating system monitors the exhaust gas temperature at the second stage low pressure turbine nozzles. The ITT indicating system also sends ITT data to the FADEC. The ITT signals are direct reading from the thermocouple probes.

4 1 0 1 5 0 0 1 1 2 7 7 D S

Operation (3.C.c) Each thermocouple lead assembly has three ITT probes of two thermocouple elements each and supplies input to the FADEC channels A and B. The ITT probes supply analog signals in relation to the ITT. The FADEC uses ITT signals for the functions that follow (engine control and indication): - Show ITT on the EICAS - Engine hot start and wet start (no ignition) logic - Engine roll back protection

EFFECTIVITY ALL

77-21-10


ENGINE INDICATING

Page 14 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

5 1 0 1 1 0 0 1 1 2 7 7 D S

TRAINING MANUAL

ITT Harness INTERTURBINE TEMPERATURE INDICATION - ITT EFFECTIVITY ALL

77-21-10


ENGINE INDICATING

Page 15 Sept 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ITT EXCEEDANCE INDICATION Identification (1.A.a) The ITT EXCEEDANCE is indicated by a change of color on the ITT dial and digital readout display of the EICAS.

6 1 0 1 5 0 0 1 1 2 7 7 D S

Purpose (2.B.b) ITT limit exceedence display requirements are complex. Limits will change as operating conditions change. This limit can change based on the following conditions: - Starts - different limits will apply for both ground and air starts. - Takeoff - Three separate take off thrust modes will trigger different ITT amber and red indications.The levels will also change on Takeoff in the event of a one engine out condition.(OEI) Takeoff limits apply for only five minutes after the takeoff thrust set then the ITT limit is set to the max continuous value. Takeoff 1 (T/O1) Takeoff 2 (T/O2) Takeoff 3 (T/O3) - Go-Around (GO) Go-around limits apply for only five minutes after the go-around thrust set then the ITT limit is set to the max continuous value. - Max Continuous (CO)

EFFECTIVITY

The MAU shall follow ITT exceedence logic for implementation of the EICAS display of ITT. This logic is based on the following inputs to the MAU: - Aircraft Speed - Critical Engine Failure Speed - ITT - ITT Normal Limit - ITT max Limit - ITT Amber Flag - Go Around Flag - Engine Running Refer to the maintenance manual for exceedence levels based on engine operating conditions. Note: If an ITT exeedance is reached before V1 minus 15 knots, the indication will be displayed on EICAS. After V1 minus 15 knots, the exceedance will be masked until landing. Exception being if the maximum of 983 C is reached which always results in an indication.

ALL

77-21-10


ENGINE INDICATING

Page 16 Sept 06

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

7 1 0 1 1 0 0 1 1 2 7 7 D S

INTERTURBINE TEMPERATURE INDICATION - ITT EFFECTIVITY ALL

77-21-10


ENGINE INDICATING

Page 17 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

VIBRATION INDICATION SYSTEM Identification (1.A.a) The engine vibration monitoring (EVM) indicating system will monitor both core and fan vibrations and supply a readout of both to the cockpit. The EVM system uses the components that follow: - No.1 bearing vibration sensor (Z1BRG) - Fan frame compressor case vibration sensor (ZFFCC) - N1 speed sensor - N2 speed sensor - EVM signal conditioner - MAU - EICAS - FADEC Purpose (2.B.b) The EVM indicating system supplies continuous data to the EICAS.

8 1 0 1 5 0 0 0 0 0 7 7 D S

Operation (3.C.c) The EVM system sends N1 RPM, N2 RPM, and vibration data to the EVM signal conditioner located in MAU 3. The EVM signal conditioner receives and amplifies the accelerometer signal which then sends an analog signal to the EICAS for monitoring. Before being displayed, the FADEC receives the sensor signal in mils and converts it to Aircraft Units. The highest value for aircraft units that can be displayed is 5. In the event of maximum display, the maintenance data computer and the flight data recorder can continue to record levels up to 10 aircraft units.

EFFECTIVITY ALL

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ENGINE INDICATING

Page 18 Mar 06

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

VIB EXCEEDENCE RESET VIB EXCEEDENCE PEAK VALUE AND TIME VIBE LIMITS N1 AND N2 MILLS TO AU CONVERSION

VIBE/SPEED SIGNALS

MAU 3

VIBE N1, N2 AND EXCEEDENCE LEVELS

9 1 0 1 1 0 0 0 0 0 7 7 D S

Z1BRG

TREND AND EXCEEDENCE DATA

ZFFCC

ACCELS, N1, N2

VIBRATION SYSTEM INTERFACE EFFECTIVITY ALL

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ENGINE INDICATING

Page 19 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ACCELEROMETERS Identification (1.A.a) There are two accelerometers. The first accelerometer is on the fan frame compressor case at the 2 o'clock position. (ZFFCC) The second accelerometer is on the No.1 bearing housing and the connector can be accessed at the 7 o'clock position on the fan frame compressor case. (Z1BRG) The ZFFCC accelerometer is an LRU and can be changed if failed. The Z1BRG accelerometer is located in the forward sump and is not an LRU. Purpose (2.B.b) The accelerometers measure the engine movement and feed the information to the vibration monitoring system. Operation (3.C.c) The accelerometers supply a small electrical output. The output level changes when the engine structure moves in the radial direction. The output difference changes in proportion to the vibration level of the engine.

0 2 0 1 5 0 0 0 0 0 2 7 D S

EFFECTIVITY ALL

72-00-00


ENGINE INDICATING

Page 20 May 07

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

1 2 0 1 1 0 0 0 0 0 2 7 D S

EFFECTIVITY ALL

72-00-00


ENGINE INDICATING

Page 21 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

AVM SIGNAL CONDITIONER Identification (1.A.a) The engine vibration monitoring (EVM) signal conditioner is in MAU 3 located in the mid avionics compartment. Purpose (2.B.b) The EVM signal conditioner has these functions: - Calculates engine vibration that shows on the EICAS - Isolates EVM system failures - Keeps historical engine vibration and system failure data in memory Operation (3.C.c) The EVM signal conditioner receives and amplifies the accelerometer signal. The EVM signal conditioner sends an analog signal to the MAU. The data is read out on EICAS.

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Training Information Points (3.E.c) No external equipment is necessary to get the data from the signal conditioner. Refer to the applicable AMMs for instructions.

EFFECTIVITY ALL

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ENGINE INDICATING

Page 22 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

MAU 3

Mid Avionics Compartment

MAU 3

EVM SIGNAL CONDITIONING EFFECTIVITY ALL

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ENGINE INDICATING

Page 23 Sept 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

OIL FILTER BYPASS/LOW OIL PS INDICATION Identification (1.A.a) Both the OIL FILTER BYPASS and LOW OIL PRESSURE indication is a crew alert message that indicates on EICAS. Purpose (2.B.b) The OIL FILTER BYPASS indication alerts the crew to a engine oil filter bypass condition. The oil filter impending bypass switch monitors the differential pressure between the lube filter inlet and lube filter discharge. The normally open switch closes when the pressure drop across the filter element rises to 21-26 psid and opens at 9 psid minimum on falling pressure. The actual bypass around the filter will begin at 41 psid. The switch contains a bimetallic interlock to prevent actuation below 100 deg F oil temperature to prevent false actuation on a cold engine start.

4 2 0 1 5 0 0 0 0 0 0 7 D S

The LOW OIL PRESSURE EICAS warning indication alerts the crew to an engine low oil pressure condition. The lube oil pressure switch measures the differential between the lube filter out pressure and the AGB internal pressure. It provides a discrete output to the aircraft. The switch is open at normal engine oil pressure and closes on falling pressure of 25 psid.

EFFECTIVITY ALL

70-00-00


ENGINE INDICATING

Page 24 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

OIL FILTER IMPENDING BYPASS SWITCH

LOW OIL PRESSURE SWITCH

5 2 0 1 1 0 0 0 0 0 0 7 D S

OIL FILTER BYPASS/LOW OIL PS INDICATION

EFFECTIVITY ALL

70-00-00


ENGINE INDICATING

Page 25 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FUEL FILTER BYPASS/LOW FUEL PS INDICATION Identification (1.A.a) Both the FUEL FILTER BYPASS and LOW FUEL RESSURE indication is a crew alert message that indicates in the cockpit on EICAS. Purpose (2.B.b) The impending fuel filter bypass switch monitors the differential pressure between the fuel filter inlet and discharge. The sensor is a normally open discrete switch and closes when the pressure drop exceeds 23-26 psid across the filter. The switch will open on 13 psid falling pressure. The low fuel pressure switch will sense the fuel pressure in the main fuel pump supply line and provides a discrete signal whenever the pressure drops below 5 psig. At this point, the aircraft boost pumps automatically come ON and remain ON until the engine is commanded OFF.

6 2 0 1 5 0 0 0 0 0 0 7 D S

EFFECTIVITY ALL

70-00-00


ENGINE INDICATING

Page 26 June 08

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

FUEL FILTER DP SWITCH

FUEL FILTER

7 2 0 1 1 0 0 0 0 0 0 7 D S

FUEL INLET PIPE TO PUMP

LOW FUEL PRESSURE SWITCH

FUEL FILTER BYPASS/LOW FUEL PS INDICATION

EFFECTIVITY ALL

70-00-00


ENGINE INDICATING

Page 27 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE OIL SYSTEM

EFFECTIVITY ALL

79-00-00


ENGINE OIL SYSTEM

Page 1 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE OIL SYSTEM Objectives Given an objective exercise, the student will identify: - The oil system of the CF34-10E engine (1.A.a) - The oil system distribution (1.A.a) - The oil storage system (1.A.a) - The location of the oil level/temperature sensor (1.A.a) - The location of lube and scavenge pump (1.A.a) - The location of the oil filter module (1.A.a) - The location of the oil filter bypass sensor (1.A.a) - The location of the oil pressure transmitter (1.A.a) - The location of the oil pressure switch (1.A.a) - The location of the electrical chip detector (1.A.a) - The sump arrangement (1.A.a) - The sump pressurization (1.A.a) - The forward sump pressurization/venting (1.A.a) - The aft sump pressurization/venting (1.A.a)

2 0 0 1 5 0 0 0 0 0 9 7 D S

Given an objective exercise, the student will select the purpose of: - The oil system of the CF34-10E engine (2.B.b) - The oil system distribution (2.B.b) - The oil storage system (2.B.b) - The oil level/temperature sensor (2.B.b) - The lube and scavenge pump (2.B.b) - The oil filter module (2.B.b) - The oil filter bypass sensor (2.B.b) - The oil pressure transmitter (2.B.b) - The oil pressure switch (2.B.b) - The electrical chip detector (2.B.b) - The sump arrangement (2.B.b)

EFFECTIVITY ALL

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ENGINE OIL SYSTEM

Page 2 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

- The sump pressurization (2.B.b) Given an objective exercise, the student will select the operation of: - The oil system distribution (3.C.c) - The oil level/temperature sensor (3.C.c) - The oil pressure transmitter (3.C.c) - The oil pressure switch (3.C.c) - The electrical chip detector (3.C.c) - The sump pressurization (3.C.c)

3 0 0 1 5 0 0 0 0 0 9 7 D S

EFFECTIVITY ALL

79-00-00


ENGINE OIL SYSTEM

Page 3 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE OIL SYSTEM Identification (1.A.a) The engine oil system has the subsystems that follow: - Storage - Distribution - Indicating Refer to figure, ENGINE OIL SYSTEM. Purpose (2.B.b) The oil system controls the flow of oil which lubricates the engine bearings and gears. The system also contains components that will give an indication to the cockpit of any oil system problems.

4 0 0 1 5 0 0 0 0 0 9 7 D S

EFFECTIVITY ALL

79-00-00


ENGINE OIL SYSTEM

Page 4 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

5 0 0 1 1 0 0 0 0 0 9 7 D S

EFFECTIVITY ALL

79-00-00


ENGINE OIL SYSTEM

Page 5 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

OIL SYSTEM DISTRIBUTION Identification (1.A.a) The oil distribution system has the systems that follow: – Supply system – Scavenge system – Sump vent system Purpose (2.B.b) The oil distribution system sends oil to lubricate the bearings and gears and to keep the bearings and gears at a satisfactory temperature. The scavenge system takes oil from the engine. The vent system allows for venting of air from the sumps and scavenge system.

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Operation (3.C.c) Supply System Oil from the tank enters the supply element of the main lubrication and scavenge pump. From this pressure element, the oil passes through the filter module, and then to the servo fuel oil heat exchanger and main fuel oil heat exchangers, which cool the oil. After leaving the heat exchangers, the oil flow divides into several circuits that lubricate the forward and aft sumps, the AGB, and the Transfer Gearbox (TGB).

The combined scavenge oil is routed past a master chip detector then into the oil tank deaerator and tank main compartment. Sump Vent System The lubrication system discharges forward sump vent air through the forward air/oil separator assembly and aft sump vent air through the aft air/oil separator into the vent tube located in the LPT shaft then to the aft vent tube and to atmosphere. Air in the scavenge return line is separated from the oil in the oil tank and vented to the forward sump. Refer to figure, OIL SYSTEM DISTRIBUTION.

Scavenge System The scavenge system includes four scavenge pump elements in the main lubrication and scavenge pump, with one element each for the AGB climb, AGB dive, forward sump, and aft sump.

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GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

OIL STORAGE SYSTEM

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Identification (1.A.a) The oil reservoir is mounted on the fan case at the 3 o’clock position (aft looking forward). The oil reservoir includes the components that are listed below. - Gravity-fill port with protective 10 x 10 wire-mesh screen. - Locking oil filler cap and a backup flapper valve in the tank to prevent oil from escaping if the cap is not sealed properly. - Gravity assisted flapper shut off valve also prevents oil tank over fill. - Sight glass for visual full level indication. - 10.0 quart (9.5 L) usable oil capacity and 14.6 US quart total oil capacity at 100% full. - Drain plug with positive locking means (safety wire). - Vortex-type deaerator to separate scavenge return air from the oil and direct it to the forward sump. - Tank pressurizing valve which opens at 5-9 psid above forward sump pressure, to ensure correct lubrication pump operation at all altitude conditions. The pressurizing valve has a small bleed hole to permit system pressure decay after engine shutdown. - Continuous oil level/temperature sensor with an oil level indication range of 10% to 106%.

EFFECTIVITY

The lubrication oil reservoir is of sufficient capacity to provide 24 hours of operation at the maximum hourly oil consumption rate of 0.4 qt/hr. The oil reservoir provides 10.0 quarts of usable oil capacity above the minimum level sensor reading and 14.6 US quart total oil capacity. The low oil warning is set at 2.4 quarts to allow 1.2 quarts useable reserve below the warning for one flight plus tolerance. Refer to figure, OIL STORAGE SYSTEM. Purpose (2.B.b) The oil storage system holds the lubrication oil for engine operation.

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GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

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CF34-10E

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OIL LEVEL/TEMPERATURE SENSOR Identification (1.A.a) The oil level/temperature sensor is mounted in the top of the oil reservoir. The sensor assembly includes an RTD element to sense oil tank temperature. Refer to figure, OIL LEVEL/TEMPERATURE SENSOR. Purpose (2.B.b) The sensor provides the level of oil within the oil tank to the aircraft, which converts the signal for display on EICAS. The level sensor reads oil level from 106% down to 10% of full capacity. Oil level sensor measures the engine lube oil temperature in the oil tank at the 10% level. The signal is sent to the aircraft for display.

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During takeoff, cruise and landing, low oil level is indicated by change in color of quantity from green to amber. There is no CAS message. During ground operation (other than takeoff and landing), low oil level in indicated by CAS message and a change in quantity from green to amber. Low level caution values are: - 4.0 quarts with engine running (N2>10) - 7.4 quarts with engine not running (N2<10)

Operation (3.C.c) Output from this sensor is directed to the cockpit for real time oil quantity status. In order for this to be accomplished the transmitter incorporates a number of reed switches. As magnets, mounted to a float assembly, move up/down past a series of switches, the reed switches close, inducing an electrical voltage through separated resistors of each switch. The total resistance as a voltage divider signal from each of the individual circuits is provided to the aircraft for display.

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CF34-10E

TRAINING MANUAL

Electrical Harness

Level Sensor

Temperature Sensor (10% level)

Oil Tank

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OIL LEVEL/TEMPERATURE SENSOR

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GE AIRCRAFT ENGINES

CF34-10E

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LUBE AND SCAVENGE PUMP Identification (1.A.a) The lube and scavenge pump is on the AGB. The lube and scavenge pump is a rotary vane type pump. The pump shaft is driven by the accessory gearbox and will provide oil flow any time the core engine is turning. The pump incorporates five pumping elements ; one supply element and four scavenge elements. Refer to figure, LUBE AND SCAVENGE PUMP. Purpose (2.B.b) The lube and scavenge pump delivers oil under pressure to the engine bearings and gears, and then recovers the oil for reuse.

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CF34-10E

TRAINING MANUAL

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GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

OIL FILTER MODULE Identification (1.A.a) The main oil filter is located in the lubrication filter module on the aft side of the AGB, schematically in the lubrication supply line between the pump and fuel/oil heat exchangers. The filter is disposable and utilizes a stainless steel mesh filtration medium. The oil filter module has an oil filtration rating of 10 microns nominal, and 15 microns absolute. It is required to remove 95% of particles larger than 10 microns (0.00039 inch), before the oil is directed to the bearings, gears and seals. At the 15 micron level the removal requirement is 99.1%. The CF34-10E lubrication supply filter is sized to have a 10-gram contamination capacity.

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The CF34-10E oil filter is disposable and thus cannot be cleaned. The filter is accessible for maintenance purposes by removing the filter bowl with a hex located at the bottom of the filter bowl to aid in the removal. Refer to figure, OIL FILTER MODULE. Purpose (2.B.b) The oil filter module filters the supply oil before the oil goes into the engine. The bypass relief valve allows full flow of the engine supply oil in the event the filter element becomes clogged or blocked.

In addition to an impending oil filter bypass switch, the filter module contains a filter bypass valve. The design of the bypass is such that no contaminant in the filter element can be released into the engine lubrication system when the flow is bypassed. The filter is not in the flow circuit when in bypass mode. The CF34-10E oil system does not contain an oil filter bypass indication, only the indication of oil filter impending bypass is sent to the aircraft MAU.

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CF34-10E

TRAINING MANUAL

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OIL FILTER BYPASS SENSOR Identification (1.A.a) The oil filter bypass sensor is mounted on the oil filter module.The oil filter bypass sensor is a normally open type pressure differential switch. Refer to figure, OIL FILTER BYPASS SENSOR. Purpose (2.B.b) The filter impending bypass switch monitors the differential pressure between the lubrication filter inlet and the lubrication filter discharge. It provides a discrete electrical output to the aircraft MAU.

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Operation (3.C.c) The normally open switch actuates (closes) when the pressure drop across the oil filter element rises to 21 to 26 psid and it deactuates (opens) at 9 psid minimum on falling pressure (note that the actual oil filter will begin bypass at 41 psid). This switch is not activated for temperatures below 100-130°F (38-54°C) to preclude false signals during cold starts.

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TRAINING MANUAL

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OIL FILTER IMPENDING BYPASS SENSOR EFFECTIVITY ALL

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GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

OIL PRESSURE TRANSMITTER Identification (1.A.a) The oil pressure transmitter is located on the oil filter module assembly, which is attached to the aft side of the AGB. Refer to figure, OIL PRESSURE TRANSMITTER. Purpose (2.B.b) The purpose of the oil pressure sensor is to provide a proportional DC output to the aircraft equivalent to the oil supply pressure to the bearing sumps.

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Operation (3.C.c) The oil pressure transmitter measures the difference in pressure between lube and scavenge pump output at the filter exit and AGB. This pressure differential is converted to an electrical signal that is sent to the aircraft. This oil pressure sensor is a piezo resistive device. A pressure signal comes from the oil supply downstream of the filter, and the reference or low pressure signal comes AGB pressure. The signal produced by the sensor is generated by resistance change within the unit. The signal is directed to the aircraft for display.

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CF34-10E

TRAINING MANUAL

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CF34-10E

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OIL PRESSURE SWITCH Identification (1.A.a) The low oil pressure switch is located on the aft side of the AGB. The low oil pressure switch is a normally closed switch. Refer to figure, OIL PRESSURE SWITCH. Purpose (2.B.b) The purpose of the low oil pressure switch is to provide a signal to the aircraft indicating and warning systems when oil pressure is low.

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Operation (3.C.c) The switch uses delta system pressure from the oil filter discharge and AGB pressure. When this pressure decreases to 25 PSID, the switch circuit will send a signal to the aircraft indicating low pressure. As oil pressure increases, the switch’s contacts open and the signal to the aircraft is no longer applied. Should oil pressure decrease due to a malfunction, the switch contacts will close and again send a signal to the aircraft.

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CF34-10E

TRAINING MANUAL

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GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ELECTRICAL CHIP DETECTOR Identification (1.A.a) The electrical chip detector is in a self-closing valve. The self-closing valve is on the aft face of the AGB. The chip detector is installed in the self-closing valve to prevent oil loss during removal for inspection and during operation if the detector is not reinstalled after inspection. A screen is also installed around the detector to collect nonmagnetic debris for inspection and identification. The detector can be easily removed for inspection without any other disassembly. Refer to figure, ELECTRICAL CHIP DETECTOR. Purpose (2.B.b) The purpose of the electrical chip detector is to trap magnetic particles that are suspended in the scavenge oil.

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Operation (3.C.c) The scavenge return oil is directed through the electrical chip detector, consisting of a powerful magnet with provisions for remotely measuring the resistance between poles and providing a warning signal, should an excess of metallic chips collect on the detector. The CMC fault message is inhibited until after a flight is complete. Indication is given when N2 is at or above idle and <66%.

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GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

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GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

SUMP ARRANGEMENT Identification (1.A.a) The CF34-10E engine has the following sumps: - Forward sump - Aft sump Forward Sump The forward sump consists of No. 1 ball bearing, No. 2 roller bearing,No. 3 roller bearing and No. 3 ball bearing. The forward sump consists of various seals and associated lubrication system. Aft Sump The aft sump consists of No. 4 roller bearing and No. 5 roller bearing. The aft sump consists of various seals and associated lubrication system.

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Purpose (2.B.b) The forward sump takes the load of the fan rotor, LPC rotor and the compressor rotor. The aft sump takes the load of the turbine section. The sumps provide confined lubrication to the bearings.

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CF34-10E

TRAINING MANUAL

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GE AIRCRAFT ENGINES

CF34-10E

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SUMP PRESSURIZATION Identification (1.A.a) The sump pressurization is done at two places: - Forward sump - Aft sump

vent tube discharges into a stationary vent tube. The stationary tube carries sump vent air to the end of the engine exhaust aft center body and discharges into the engine exhaust stream.

Purpose (2.B.b) The sump pressurization is done to prevent oil leakage From the sumps. Operation (3.C.c) The sump pressurization system is supplied by booster discharge air. The forward sump seals are supplied through multiple tubes and internal stationary passages. The aft sump is supplied with air that passes first through four radial stationary tubes in the middle of the front sump and then aft through a rotating annulus bounded by a duct in the bore of the HP rotor and the LPT shaft. HP rotor bore cooling air is separated from the aft sump pressurization air by the HP rotor bore duct. 6 2 0 1 5 0 0 0 0 0 9 7 D S

All sump oil and air seals are multi-tooth labyrinth designs. Each sump has four air-oil seals. Where required, pressurization air is separated from other secondary air flows by labyrinth air-air seals. Forward and aft sumps are each vented through LPT shaftmounted dynamic air-oil separators. Both air-oil separators discharge into a central rotating vent tube mounted inside the LPT shaft. The rotating

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GE AIRCRAFT ENGINES

CF34-10E

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FORWARD SUMP PRESSURIZATION/VENTING Identification (1.A.a) The CF34-10E forward sump has labyrinth oil seals located at the forward No. 1 ball bearing air-oil seal and at the aft end, No. 3 radial bearing air-oil seal.

empties into the rotating center vent tube, located within the LPT shaft. This rotating center vent tube leads to an extended stationary center vent tube that discharges through the aft end of the engine.

Purpose (2.B.b) The forward sump pressurization/venting system helps in preventing oil leakage from the forward sump.

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Operation (3.C.c) Pressurization air is supplied to the No. 1 bearing air-oil seal through stationary tubes. Pressurization air (and HP rotor bore cooling air) is supplied to the No. 3 bearing aft air-oil seal through cast-in passages within the seal support. The pressurization chamber outside the No. 3 bearing aft air-oil seals are provided with drains to route any oil seal leakage overboard. Inside the forward sump, a pair of seals, (No. 2 roller bearing air-oil and No. 3 ball bearing forward air-oil seal) share a single support and seal against the LPT and HPC shafts, respectively. These two seals are pressurized through four radial stationary tubes which also are the source of pressurization air for the aft sump. The forward sump air-oil separator is a multi-passage radial flow dynamic device attached to the fan forward shaft. As vent air flows through the separator passages, entrained oil is centrifuged back into the sump. Vent air

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CF34-10E

TRAINING MANUAL

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GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

AFT SUMP PRESSURIZATION/VENTING Identification (1.A.a) The CF34-10E aft sump has two oil seals, (No. 4 roller bearing inner air-oil seal and No. 4 roller bearing outer air-oil seal) between the HP and LP rotors near the No. 4 roller bearing. Purpose (2.B.b) The aft sump pressurization/venting system helps in preventing oil leakage from the aft sump.

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Operation (3.C.c) Pressurization air arrives to these seals via the annular passage bounded by the HP rotor bore duct and LPT shaft. The No. 5 roller bearing side of the sump also has two oil seals, No. 5 roller outer air-oil seal and No. 5 roller inner air-oil seal. Pressurization air for these seals is passed through holes in the LPT shaft flange from the No. 4 bearing side of the sump. The pressurization circuit surrounds the aft sump helping to maintain moderate temperatures within the sump. Pressurized air-to-air seals, (No. 4 roller inner air-air seal, and No. 5 roller outer aft air-air seal) separate the rotor bore cooling flow from the sump pressurization circuit. The No. 5 roller outer aft air-air seal acts with the No. 5 roller outer forward air-air seal to form an air vent and oil drain chamber that is connected to the aft center body volume by 4 passages through the No. 5 roller bearing support. Potential aft sump oil seal leakage is passed through this vent/drain chamber, into the aft center body which has a

EFFECTIVITY

drain hole open to the engine exhaust stream. The aft sump air-oil separator is a multi-passage radial flow dynamic device attached to the LPT shaft. As vent air flows through the separator passages, entrained oil is centrifuged back into the sump. Vent air empties into the rotating center vent tube, located within the LPT shaft. This rotating center vent tube leads to an extended stationary center vent tube that discharges through the aft end of the engine.

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CF34-10E

TRAINING MANUAL

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Page 31 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE STARTING

EFFECTIVITY ALL

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ENGINE STARTING

Page 1 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE STARTING Objectives Given an objective exercise, the student will identify: - The CF34-10E engine starting system (1.A.a) - The distribution and control (1.A.a) - Selected flight compartment switches (1.A.a) - The location of the engine start valve (1.A.a) - The location of the air turbine starter (1.A.a) - Selected components of the engine starting indication (1.A.a) Given an objective exercise, the student will identify the purpose of: - The CF34-10E engine starting system (2.B.b) - Selected flight compartment switches (2.B.b) - The engine start valve (2.B.b) - The air turbine starter (2.B.b) - Selected components of the engine starting indication (2.B.b)

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Given an objective exercise, the student will identify the operation of: - The CF34-10E engine starting system (3.C.c) - The distribution and control (3.C.c) - The engine start valve (3.C.c) - The air turbine starter (3.C.c)

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GE AIRCRAFT ENGINES

CF34-10E

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ENGINE STARTING

Page 3 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

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ENGINE STARTING SYSTEM Identification (1.A.a) The engine starting system contains an air turbine starter (ATS) and a starter control valve for each engine. The starting system also utilizes the engine full authority digital engine control (FADEC), ignition system and fuel system as well as the aircraft APU, EICAS, and electrical systems.

commanded by the FADEC. The aircraft also manages the bleed system interface during starts.

The engine incorporates a pneumatic starter and is designed to consistently make satisfactory ground and starter-assisted air starts within the Mach number, altitude, ambient temperature, and power extraction limits. Maximum aircraft accessory loads, referenced to the gas generator rotor, are permitted during starterassisted starting. No wing anti-ice or ECS bleed extraction is permitted during the start.

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Purpose (2.B.b) The purpose of the engine starting system is to provide means for the engine to obtain sufficient rotor speed to initiate combustion light-off and obtain self-sustaining engine propulsion. Operation (3.C.c) Engine-starting is a combined aircraft and FADEC operation. The FADEC controls fuel flow, the starter command, and the ignition command. The aircraft controls the starter valve, the engine-driven pump (EDP), and switches power to the ignition exciters as

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Page 4 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

ENGINE 1 IGNITION A IGN A IGN B

ENGINE 1 ENGINE 2 EXCITER 1A RLY START

ENGINE 1

ENGINE 1-2 EXCITER A CMD

ENGINE 1 AND 2 IGNITION OVERRIDE

ENGINE 1 AND 2 FADEC CHANNEL A IGNITION

ASCB

OVRD

OVRD

ENGINE 1 FADEC CHANNEL B IGNITION

ENGINE 1 IGNITIO N B

ENGINE 2

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ENGINE 2 FADEC CHANNEL B IGNITION

IGN B IGN A

ENGINE 2 IGNITION A ENGINE 2 IGNITION B ENGINE 2 START VALVE ENGINE 1 START VALVE

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Page 5 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

DISTRIBUTION AND CONTROL Identification (1.A.a) The starter control is an integrated aircraft/FADEC function. The FADEC controls the starter cutout and start abort. Operation (2.B.b) Once fuel has been delivered from the aircraft to the engine, engine fuel control is performed exclusively by the FADEC. The pilot commands fuel on by moving the Stop/Run/Start Switch from the Stop position to the Start position. If the pilot has requested fuel on via the Stop/Run/Start Switch, the FADEC turns fuel on as per the following procedure:

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1. On the ground, the FADEC will open the metering valve when N2 is greater than 19.1% (of 18018 RPM). The FADEC will close the metering valve under the following conditions: (a) Hot start - FMV and ignition is switched switched off when ITT is greater than 740 deg C before ground idle is reached. (b) Hung start - FMV and ignition ignition off when engine has has light off but N2 dot goes near zero before ground idle is reached. (c) Rollback on ground - FMV and ignition off ifif engine N2 rolls back below 55%.

EFFECTIVITY

2. For in-flight starter assisted starts, the FADEC will open the metering valve when N2 is greater than approximately 19.1% (of 18018 RPM). If N2 has not reached the 19.1% speed after 15 seconds, the metering valve will be opened. 3. For in-flight windmill starts, the FADEC will open the metering valve when N2 is greater than approximately 6.9% (of 18018 RPM) speed. If N2 has not reached 6.9% speed after 15 seconds, the metering valve will be opened. 4. Independent of flight or ground, the FADEC will shut the engine down under the following conditions: (a) Loss of two fuel metering valve feedback signals (b) Large WF feedback soft fault (not a commanded shutdown) (c) Loss of four N2 signals (d) Three over speed trips within 30 seconds or one over speed trip below 20,000 ft. The 20,000 ft level is to help prevent controllability issues during an approach. The take-off ceiling for the 190 is 15,000 ft.

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Page 6 June 08

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL


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ENGINE STARTING

Page 7 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

The pilot commands fuel shutoff by moving the Stop/ Run/Start Switch to the Stop position. If Stop is commanded in-flight, the metering valve is closed and a momentary FADEC hardware reset is commanded. On the other hand, if Stop is commanded on the ground, the FADEC shuts down the engine by commanding a test of the over speed system. This test will be delayed by at least 0.2 seconds to allow the aircraft electrical system to transfer electrical loads from engine-supplied power to aircraft-supplied power.

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Page 9 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

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FLIGHT COMPARTMENT SWITCHES Identification (1.A.a) The start system interfaces with a three position ignition switch (OFF/AUTO/OVRD) that allows the pilot to send ignition commands to the FADEC and by a three position run switch (STOP/RUN/START) that will initiate the start or stop sequence. Operation (2.B.b) The start sequence is initiated as follows: - The Start/Stop Start/Stop switch is set set to Stop - The Ignition Selector Selector Switch Switch is set to Auto Auto - The Stop/Start Stop/Start Switch is is set to Start. Start. This is a momentary switch. - The Start switch switch signal is sent to the FADEC FADEC - The FADEC sends an Energize Energize Starter Starter command signal and opens the starter SOV.

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Starter rotation continues until any of the following occur: - The pilot aborts aborts the start start by selecting selecting STOP on the selector switch. - The FADEC closes closes the Starter Starter SOV when N2 N2 starter cutout speed is reached. - The FADEC closes closes the starter starter SOV if the Fire Fire Handle is pulled. Ignition switch positions can be for the following functions: AUTO - Normal start start position. The FADEC will alternate

EFFECTIVITY

between A and B ignition on each start in order to reveal any igniter or exciter faults. Both igniters will come ON in AUTO position if the following occurs: - If an aircraft stall condition condition is indicated. indicated. - If engine detects a flame out - When the engine engine is running running or in-flight start start - A missed light off off is detected detected (ground (ground only) - An igniter fault fault is detected detected while performing performing a ground start - Loss of of ARINC busse busses s OVRD – Commands Commands ignition ignition A and B “ON” through FADEC and by directly powering the relay coils located in the Emergency Integrated Control Center (EICC) based on condition. Which igniter is command “ON” is PS3.0 dependent. OFF - Off position position can selected for for the following conditions: - Commands engine engine shutdown. shutdown. When ignition is commanded OFF, fuel will also be commanded off. - Fire handle handle pulle pulled d - Hot start start - Hung Hung start start - Rollback Rollback on ground ground - Starting cycle cycle complete (engine is above above starter cutout speed)

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Page 10 Dec 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

IGNITION SWITCH PANEL

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ENGINE STARTING

Page 11 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

The START/STOP switch positions can be for the following functions: START - This position initiates the start sequence. It is a momentary switch position. RUN - This position is the normal run condition. The momentary start switch returns to this position after start sequence is initiated and remains there until a command to stop engine is done. STOP - With the TLA at the Idle position, fuel shutoff is accomplished by moving the Start/Stop selector switch to the STOP position. Actual fuel shutoff is accomplished by the FADEC closing the fuel metering valve and energizing the over speed shutoff solenoid.

2 1 0 1 5 0 0 0 0 0 0 8 D S

EFFECTIVITY ALL

80-00-00


ENGINE STARTING

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EFFECTIVITY ALL

80-00-00


ENGINE STARTING

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ENGINE START VALVE Identification (1.A.a) The starter control valve (SCV) is a spring-loaded, closed, pneumatically actuated, electrically controlled, butterfly type shutoff valve with an open-position indication switch. The valve incorporates a visual position indicator and a manual override feature for operating the unit with loss of electrical power. The manual override feature is a 3/8” square internal drive.

CF34-10E

TRAINING MANUAL

The starter control valve is controlled by the FADEC using an aircraft powered 28 VDC solenoid. The twoposition valve is normally spring-loaded and air pressure closed. When the solenoid is energized, starter duct pressure is used to open the starter control valve. A manual override “socket drive” is provided to open the valve in case of physical or electrical failure.

Operational air can come from a ground power unit (GPU), an auxiliary power unit (APU), or another operating engine. Refer to figure, ENGINE START VALVE. Purpose (2.B.b) The starter control valve controls the air flow to the starter. 4 1 0 1 5 0 0 0 0 0 0 8 D S

Operation (3.C.c) The valve controls the rise rate of inlet pressure to the engine starter when energized open. The valve moves from full closed to full open in 4 to 8 seconds. The normal operating range for the air inlet pressure is 25 to 48 psig. When deenergized, the valve closes, under spring force, to shut off airflow to the starter. The valve moves from full open to full closed in 2 to 4 seconds.

EFFECTIVITY ALL

80-00-00


ENGINE STARTING

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TRAINING MANUAL

5 1 0 1 1 0 0 0 0 0 0 8 D S

EFFECTIVITY ALL

80-00-00


ENGINE STARTING

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TRAINING MANUAL

AIR TURBINE STARTER Identification (1.A.a) The starter is a single-stage air turbine clamped to the AGB forward face at the 9:00 adapter pad by a Vcoupling clamp. A locator pin is provided between the mounting flange interfaces to accurately position the starter to the engine. Refer to figure, AIR TURBINE STARTER. Purpose (2.B.b) The starter is used to accelerate the engine core from 0% N2 to self-sustaining RPM and provides wet/dry motoring during maintenance practices.

6 1 0 1 5 0 0 0 0 0 0 8 D S

Operation (3.C.c) The starter is controlled by the FADEC-triggered starter control valve. When the FADEC signals the shutoff valve open, airflow passes through the starter turbine, creating rotation. The starter output shaft turns the accessory gearbox drive train and rotates the core engine. A Sprag-type overrunning clutch automatically disconnects the turbine from the output shaft after the SCV is closed and the starter turbine slows.

EFFECTIVITY ALL

80-00-00


ENGINE STARTING

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TRAINING MANUAL

7 1 0 1 1 0 0 0 0 0 0 8 D S

EFFECTIVITY ALL

80-00-00


ENGINE STARTING

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CF34-10E

TRAINING MANUAL

ENGINE EXHAUST


78-00-00 ENGINE EXHAUST

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TRAINING MANUAL

ENGINE EXHAUST Objectives Given an objective exercise, the student will identify: - The CF34-10E engine engine exhaust exhaust system (1.A.a) (1.A.a) - The primary exhaust exhaust system system (1.A.a) - The location of the thrust reverser reverser (1.A.a) - The location of the synchronized synchronized locking actuators actuators (SLA) (1.A.a) - The location of the synchronized synchronized feedback feedback actuators (SFA) (1.A.a) (1.A.a) - The location of the isolation isolation control unit unit (ICU) (1.A.a) - The location of the direction direction control control unit (DCU) (1.A.a) - The location of of the cowl lock (CL) (CL) (1.A.a) - The location of the synchronizin synchronizing g shafts (1.A.a) (1.A.a) - The location of the sync sync tube assembly assembly (1.A.a) - The location of the manual drive assembly assembly (1.A.a) - The location of the flex tube assembly assembly (1.A.a) - The location of the ca cascade scade vanes vanes (1.A.a) - The location of of the translating translating cowls (1.A.a) (1.A.a) - The location of the thrust reverser reverser opening actuators actuators (1.A.a) - The location of the ground maintenance maintenance override override switch (1.A.a) - The thrust reverser reverser deenergized deenergized stage (1.A.a) (1.A.a) - The thrust reverser reverser system system indication (1.A.a) (1.A.a) 2 0 0 1 5 0 0 0 0 0 8 7 D S

Given an objective exercise, exercise, the student will select the purpose of: - The CF3410 engine engine exhaust system system (2.B.b) - The primary exhaust exhaust system system (2.B.b) - The thrust thrust reverser reverser (2.B.b) (2.B.b) - The synchronized synchronized locking actuators actuators (SLA) (2.B.b) (2.B.b) - The synchronized synchronized feedback actuators actuators (SFA) (2.B.b) - The isolation control unit unit (ICU) (2.B.b) (2.B.b) - The direction control control unit unit (DCU) (2.B.b) - The cowl lock (CL) (2.B.b)



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TRAINING MANUAL

- The synchronizing synchronizing shafts shafts (2.B.b) (2.B.b) - The sync tube assembly assembly (2.B.b) (2.B.b) - The manual manual drive assembly assembly (2.B.b) (2.B.b) - The flex tube tube assembly assembly (2.B.b) (2.B.b) - The cascade cascade vanes vanes (2.B.b) (2.B.b) - The translating translating cowls cowls (2.B.b) (2.B.b) - The thrust reverser reverser opening opening actuators (2.B.b) (2.B.b) - The ground maintenance maintenance override override switch (1.A.a) (1.A.a) - The thrust thrust reverser reverser deployed deployed - stage 1 (2.B.b) (2.B.b) - The thrust thrust reverser reverser deployed deployed - stage 2 (2.B.b) (2.B.b) - The thrust thrust reverser reverser deployed deployed - stage 3 (2.B.b) (2.B.b) - The thrust reverser reverser stow stow stage (2.B.b) (2.B.b) - The thrust reverser reverser system system indication (2.B.b) (2.B.b)

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Given an objective exercise, the student will select the operation of: - The primary exhaust exhaust system (3.C.c) - The isolation control unit unit (ICU) (3.C.c) (3.C.c) - The direction direction control unit (DCU) (3.C.c) - The cowl cowl lock (CL) (3.C.c) (3.C.c) - The thrust reverser reverser opening opening actuators (3.C.c) (3.C.c) - The ground maintenance maintenance override override switch (3.C.c) (3.C.c) - The thrust reverser reverser de-energized de-energized stage (3.C.c) (3.C.c) - The thrust thrust reverser reverser deployed deployed - stage 1 (3.C.c) (3.C.c) - The thrust thrust reverser reverser deployed deployed - stage 2 (3.C.c) (3.C.c) - The thrust thrust reverser reverser deployed deployed - stage 3 (3.C.c) (3.C.c) - The thrust reverser reverser stow stow stage (3.C.c) (3.C.c) - The thrust reverser reverser system system indication (3.C.c) (3.C.c) Given an objective exercise, the student will select the maintenance tip of: - The thrust reverser reverser deenergized deenergized stage (4.D.d) (4.D.d)



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TRAINING MANUAL

ENGINE EXHAUST SYSTEM Identification (1.A.a) The engine exhaust system consists of two main subsystems: - Primary exhaust system - Thrust Thrust reverser reverser system Purpose (2.A.a) The engine exhaust system of the CF34-10E engine provides the mechanism to discharge the air from the propulsion system. Refer to figure, ENGINE EXHAUST SYSTEM.

4 0 0 1 5 0 0 0 0 0 8 7 D S



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TRAINING MANUAL

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PRIMARY EXHAUST SYSTEM

assembly at the 12 o’clock position.

Identification (1.A.a) The primary exhaust system consists of a chevron nozzle, centerbody, and a center vent tube.

Centerbody There is a forward centerbody and a aft centerbody for the CF34-10E engine.

Refer to figure, PRIMARY EXHAUST SYSTEM.

Center Vent Tube The center vent tube is at the center of the primary exhaust system.

The following design considerations have been incorporated into the CF34-10E propulsion system exhaust: - All parts of the primary primary exhaust system system are made of Titanium, fireproof and are designed to to withstand the anticipated operating environment. - The propulsion propulsion system drains drains and vents vents are positioned forward of the exhaust gases of the primary exhaust system. - Provisions for core core compartment compartment and sump ventilation have been provided in the exhaust system. In addition, the centerbody incorporates a vent tube to discharge sump air into the flow. 6 0 0 1 5 0 0 0 0 1 8 7 D S

CF34-10E

Chevron Nozzle The chevron (primary exhaust) nozzle assembly is composed of a forward flange for attachment to the engine turbine frame outer flange and a conical section sheet metal skin welded welded to the forward flange. A fire shield is also incorporated into the chevron nozzle


Purpose (2.B.b) The primary exhaust system provides a fixed area annulus for exhausting the core engine gas stream flow and provides a continuation of the aerodynamic cowling from the aft core cowl interface. The primary exhaust nozzle is of a chevron design to reduce the engine core noise levels. The purpose of the centerbody is to direct the core exhaust flow. Operation (3.C.c) Primary air is that air which enters the engine near the fan blade platform, continues through the booster compressor, high pressure compressor, the combustor, the high and low pressure turbines, and then is accelerated and exhausted to atmosphere through the primary nozzle. The inner wall of the primary nozzle and the outer wall of the centerbody form the primary nozzle flowpath.


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THRUST REVERSER Identification (1.A.a) The thrust reverser is constructed in two halves that are hinged to the aircraft pylon at the top and latched together at the bottom to permit opening the thrust reverser for engine access or removal. The thrust reverser consists of: - Translating cowls - Cascades - Structural components - Thrust reverser actuation system

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Purpose (2.B.b) The thrust reverser assembly forms the fan air stream exhaust nozzle when stowed and reverses the direction of the fan air stream when deployed.

The thrust reverser actuation system consists of the following major components: - Two synchronized locking actuators (SLA) - Manual drive assembly (2) - Two synchronized feedback actuators (non-locking) (SFA) - One isolation control unit (ICU) - One direction control unit (DCU) - One cowl lock (CL) - One set of (3) synchronizing shafts - Sync tube assembly (3) - Flex tube assembly - Ground Maintenance Override Switch (GMO) Refer to figure, THRUST REVERSER.



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TRAINING MANUAL

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TRAINING MANUAL

SYNCHRONIZED LOCKING ACTUATORS (SLA) Identification (1.A.a) There are two locking actuators. The locking actuators are attached to the torque box and to the translating cowl. No. 1 and No. 2 actuators consist of piston and nut assembly, screw shaft, tine-lock mechanisms with a manual unlock facility, worm shaft/worm wheel gearing, control microswitch, and monitoring microswitch. All components are in the actuator body. Actuator is gimbaled to allow relative movement. The synchronized locking actuator (SLA) has an internal lock mechanism consisting of three radially located locking keys, captured in slots in the cylinder, and a lock sleeve. The lock can be released by a manual unlock handle. Both SLA have a simplex lock switch, which provides lock status information. 0 1 0 1 5 0 0 0 0 0 8 7 D S

Refer to figure, SYNCHRONIZED LOCKING ACTUATORS. Purpose (2.B.b) The locking actuators move the translating cowl, mechanically lock it in the stowed position, and give a locked indication to the FADEC.



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TRAINING MANUAL

1 1 0 1 1 0 0 0 0 0 8 7 D S



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TRAINING MANUAL

MANUAL DRIVE ASSEMBLY Identification (1.A.a) A manual drive is built into the head side of each locking actuator. The drive allows manual operation of the thrust reverser. It has a torque limiter mechanism and is driven using a 3/8” square driver. Refer to figure, MANUAL DRIVE ASSEMBLY. Purpose (2.B.b) The drive deploys and stows the thrust reverser translating cowl during maintenance. It limits the torque from the manual drive input to prevent damage to the synchronizing system. Maintenance Note: Both locking actuators must be unlocked before manual deployment of the translating cowl through the left or right manual drive assembly. 4 2 0 1 5 0 0 0 0 0 8 7 D S



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TRAINING MANUAL

SYNCHRONIZED FEEDBACK ACTUATORS (SFA) Identification (1.A.a) There are two feedback actuators. The feedback actuators are attached to the torque box and to the translating cowl. The feedback actuators consists of: - Piston and nut assembly - Screw shaft - Linear variable differential transducer (LVDT) - Worm shaft/worm wheel gearing All components are in the actuator. Actuator is gimbaled to allow relative movement. The synchronized feedback actuator (SFA) has no internal lock, but is in design and construction similar to the locking actuator. The SFA has a feedback mechanism consisting of a LVDT actuated by a rod attached to the lead screw. The LVDT provides an indication of the actuator extension. 2 1 0 1 5 0 0 0 0 0 8 7 D S

Refer to figure, SYNCHRONIZED FEEDBACK ACTUATORS. Purpose (2.B.b) The feedback actuators move the translating cowl and give position indication to the FADEC.



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TRAINING MANUAL

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TRAINING MANUAL

ISOLATION CONTROL UNIT (ICU) Identification (1.A.a) The ICU is located in the engine pylon behind an access panel on the right hand side. The ICU consists of isolation solenoid valve, cone type isolation valve, manual inhibit lever, pressure switch, and filter. The ICU is powered through a FADEC ground relay.

positions a stop for the valve and may be locked in both states by a pip pin. Manual inhibit is indicated by a single pole switch. The valve should also be locked by the inhibit lever in the closed position for aircraft dispatch with the reverser locked out.

Refer to figure, ISOLATION CONTROL UNIT. Purpose (2.B.b) The ICU isolates the thrust reverser system from the aircraft hydraulic supply when the thrust reverser is not in use. The ICU provides hydraulic power to the DCU during thrust reverser operation.

4 1 0 1 5 0 0 0 0 0 8 7 D S

Operation (3.C.c) The ICU contains the isolation valve, which is normally held in closed position by a spring and by the return pressure. The isolation valve is moved by energizing the normally closed simplex solenoid valve. This directs pressure to the pilot area of the spool, shuttling the valve so as to allow pressure to the cowl lock and the rod side of the actuators. Main pressure passes through a cartridge inlet screen. A simplex pressure switch monitors downstream pressure. The valve may be locked in the closed position for engine maintenance by an inhibit lever. The lever



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TRAINING MANUAL

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TRAINING MANUAL

DIRECTION CONTROL UNIT (DCU) Identification (1.A.a) The DCU is on the aft side of the left hand torque box, between the feedback and locking actuators, downstream of the ICU. The DCU is an electrically controlled hydraulic directional control valve which is powered through a FADEC ground relay. Refer to figure, DIRECTION CONTROL UNIT. Purpose (2.B.b) The DCU controls the direction of the TRAS through the feedback and locking actuators. Operation (3.C.c) The DCU contains the direction control valve, which has two positions corresponding to stow and deploy. The valve is normally held in the stow position by a spring and by the return pressure. 6 1 0 1 5 0 0 0 0 0 8 7 D S

The DCU incorporates a normally closed, single coil solenoid valve with integral electrical connector. When energized (after delivery of pressure from the ICU), the solenoid valve connects pressure to the head side of the actuators unlocking the actuator locks. Note, the actuator locks unlock at a much lower pressure even before any actuator movement.



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TRAINING MANUAL

COWL LOCK (CL) Identification (1.A.a) The cowl lock (CL) is attached to the LH aft portion of the hinge beam on the thrust reverser, under a removable fairing. The CL is a hydraulically operated hook and is powered through the SPDA. Refer to figure, COWL LOCK. Purpose (2.B.b) The cowl lock prevents uncommanded deployment of the translating cowl.

8 1 0 1 5 0 0 0 0 0 8 7 D S

Operation (3.C.c) The cowl lock: - Holds the cowl in the stowed position following failure of both actuator lock mechanisms - Stays in the unlocked position while the thrust reverser is cycled - Returns the indicating switch signal to the A/C when the cowl lock has been relocked - Can be manually unlocked and pinned to allow manual movement of the translating cowls for maintenance purposes.



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TRAINING MANUAL

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TRAINING MANUAL

SYNCHRONIZING SHAFTS Identification (1.A.a) There are three synchronizing shafts for each engine. The synchronizing shafts are located inside two sync tubes and one flex tube and are connected to the worm shaft of each feedback and locking actuator. The drive shafts, which make up the transmission, are made from layers of stainless steel wires. Solid square adapters on each end of the shafts interface with the actuator worm shafts. Refer to figure, SYNCHRONIZING SHAFTS. Purpose (2.B.b) A flexible sync shaft running between the actuators synchronizes the actuator movements to help prevent unbalanced translating cowl loads and to ensure equal deployment of the two translating cowl halves.

0 2 0 1 5 0 0 0 0 0 8 7 D S



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TRAINING MANUAL

SYNC TUBE ASSEMBLY Identification (1.A.a) There are two sync tube assembly in each engine. The sync tubes are located between the feedback actuator and the locking actuators. The sync tubes are metal tubes. Purpose (2.B.b) The sync tube contains the flex shafts to the head of each actuator. This allows for mechanical synchronization between the feedback and locking actuators. The sync tubes not only contain the flex shafts but also the hydraulic pressure supplied to the “deploy” side of the actuators.

2 2 0 1 5 0 0 0 0 0 8 7 D S



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3 2 0 1 1 0 0 0 0 0 8 7 D S



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FLEX TUBE ASSEMBLY Identification (1.A.a) There is one flex tube on each engine. The flex tube is between the two feedback actuators over the top of the engine. The flex tube helps in opening the thrust reverser half’s without removing any components of the actuation system. Purpose (2.B.b) The flex tube carries a flex shaft between the two feedback actuators providing mechanical synchronization and hydraulic pressure supplied to the “deploy” side of the actuators.

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Flex Tube

7 2 0 1 1 0 0 0 0 0 8 7 D S

Left Hand TR Assembly

FLEX TUBE ASEMBLY EFFECTIVITY ALL GE PROPRIETARY INFORMATION


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CASCADE VANES Identification (1.A.a) The cascade vanes are on the torque box frames and cascade support rings on the left and right fixed structures. The vanes are composite structures. Refer to figure, CASCADE VANES. Purpose (2.B.b) When the thrust reverser is deployed, the vanes redirect the fixed fan airflow outward and forward to produce a decelerating effect during landing.

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TRANSLATING COWLS Identification (1.A.a) There are two translating cowls on each engine. The translating cowls are on the upper hinge beam and the lower latch beam assemblies. Each translating cowl is attached to its fixed structure half by an upper and a lower rail and the actuators. Two thrust reverser actuators are connected to each translating cowl by means of light alloy actuator fittings fastened to the inner bondment. The outer upper and lower edges of the translating cowl outer skin are retained by keeper fittings, which slide along the catcher tracks of the beams. A removable forward lip is fastened to the inner edge of the front frame to permit installation of the translating cowl on to the fixed structure.

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The two translating cowls are connected together by a set of two latches at the 6 o'clock location, visible from the outside. An access door is installed at 6 o'clock to gain access to the one fixed structure mid-latch, hidden behind the translating cowl bottom section. Refer to figure, TRANSLATING COWLS. Purpose (2.B.b) When in the stowed position, the translating cowl completes the smooth aerodynamic flow paths. When it is deployed, it blocks fan discharge air and directs it forward.



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THRUST REVERSER POWER DOOR OPENERS (PDO) Identification (1.A.a) The thrust reverser power door openers are at the 3 o'clock and at the 9 o'clock position on the thrust reverser frame. The PDO can be locked in one of two positions and function as a hold open rod to maintain each fan reverser half in the open position. Purpose (2.B.b) The PDO actuators help in opening the thrust reverser half for ground maintenance and inspection of the engine and accessories. Operation (3.C.c) The opening actuator is actuated by a hand pump and a flexible pipe, which are not parts of the nacelle. The actuators can be latched in two positions. 20 deg or 43 deg at it's full open position.

2 3 0 1 5 0 0 0 0 0 8 7 D S

Maintenance Note: Attempting to rapidly open the thrust reverser through the hydraulic pump may cause the pressure relief valve located next to the Hanson fitting to release to atmosphere. Wear eye protection and open slowly.



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PDO Actuator

FLA

“Hanson” Fitting for Hydraulic Hand Pump

Thrust Reverser Torque Box Engine Attach Point

Relief Valve

Reverser Attach Point

THRUST REVERSER OPENING ACTUATORS (PDO)



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GROUND MAINTENANCE OVERRIDE SWITCH (GMO) Identification (1.A.a) The ground maintenance override switch is a contact switch located on the inlet bulkhead at the 3:00 position. Refer to figure, GROUND MAINTENANCE OVERRIDE SWITCH. Purpose (2.B.b) The ground maintenance override switch allows maintenance personnel to operate the thrust reverser on the ground without the engine running. Operation (3.C.c) The ground maintenance override is a momentary switch that causes the FADEC to close both the ICU and DCU ground relays. The switch needs contact with only one FADEC channel but is interlocked with engine speed and aircraft speed. N2 must be <9.96% and aircraft speed <40 knots. Switch closure overrides the “engine not running” FADEC requirement that would normally inhibit deployment on the ground. The switch must be held in order for ground deployment to be accomplished. If the switch is closed for more than two minutes, it will be disabled. This is a safety feature that ensures that there is at least one person at the engine during ground operation. It will be enabled again after the switch is released for 30 seconds. GMO is independent of WOW to allow for operation of the TRAS while the aircraft is on jacks.



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A

Channel A&B GMO Switch

Fan Case

A

Mounted to Nose Inlet Cowl

GROUND MAINTENANCE OVERRIDE SWITCH (GMO) EFFECTIVITY ALL GE PROPRIETARY INFORMATION


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THRUST REVERSER DEENERGIZED STAGE Identification (1.A.a) In the deenergized state the actuators are at rest in the flush, stowed position. All electrical signals are removed and all locks are engaged. Clearance in the cowl lock ensures it is unloaded during normal operation. The ICU isolates the hydraulic supply such that all system components are at return pressure. Any leakages through the ICU are dissipated to return such that the thrust reverser actuation system (TRAS) cannot pressurize.

4 3 0 1 5 0 0 0 0 0 8 7 D S

Maintenance Tip (4.D.d) For manual opening of the TRAS a maintenance lever on the ICU locks the isolation control valve in the closed position. A single micro switch indicates locked position of the inhibit lever which can be locked in either power or inhibited positions by a pip pin. The cowls are over stowed using the manual drives located on the end of the synchronization shaft (mounted on the lower/locking actuators). The cowl lock hook is released by pushing upward on the hook and pinning it in the open position. The actuator internal locks may be manually disengaged by rotating the manual unlock shaft, the status of which is indicated by a limit switch. The extend and retract actuator areas are now connected to return permitting movement of the cowls from the manual drive. Note, in order to move the actuators there must be free flow of fluid to/from the aircraft return line.



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THRUST REVERSER DEPLOYED STAGE1 Purpose (2.B.b) The transcowls are over stowed to remove the pressure on the cowl locks. Operation (3.C.c) The ICU solenoid is energized directly from FADEC commanded ICU Ground Relay when the Reverser is "enabled". When the pilot commands deploy, the Thrust Control Quadrant (TCQ) switches are closed and simultaneous electrical signals are sent to release the cowl lock and to operate the DCU solenoid valve. Movement of the isolation solenoid valve initiates three actions: 1. Pressure is allowed to and through the DCU 2. Pressure is applied to the retract area of all actuators 3. A “pressure on” signal is generated by the ICU pressure switch 6 3 0 1 5 0 0 0 0 0 8 7 D S

When pressure is applied to the rod end of the actuators this causes them to retract to the over stow position where the actuator internal locks may be released. A time delay in the FADEC prevents the DCU from being energized to allow the cowl lock to get out of the way and prevent a deployment against a closed lock.



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CF34-10E

TRAINING MANUAL

THRUST REVERSER DEPLOYED STAGE-2 Purpose (2.B.b) Translating cowl deployment. Operation (3.C.c) Movement of the direction control unit spool valve permits pressure to the head areas of the actuators. As pressure increases in the actuator head chambers the internal lock piston is pulled back against it’s spring and return pressure. The spring and lock piston areas are configured such that the internal lock is released prior to actuator movement in the extend direction. At this time, supply pressure (3000 psi) is present on both sides of the actuator piston. However, due to the differential area across the actuator piston, the greater force on the head side causes the actuators to extend to full deploy position.

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The nut attached to the hydraulic piston means that the linear actuator movement causes the internal lead screw to rotate. This in turn drives the worm wheel in the actuator head. The worm wheel in each actuator is connected to its neighbor via flexible cables running within the extend hydraulic pipes. The efficiencies of the screw, worm gears and shafts act to retard a leading actuators. In this way all actuators are restrained to move in unison. As the cowls approach full deploy the actuator's snubbers engage and flow from the rod chambers is restricted to decelerate the actuators and therefore minimize impact loads.



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GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

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Page 41 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

THRUST REVERSER DEPLOYED STAGE-3 Purpose (2.B.b) The translating cowls in the deployed stage till the stow command is initiated. Operation (3.C.c) While deployed, both the ICU and the cowl lock solenoids remain powered. If an uncommanded stow is detected by the FADEC, power is automatically removed from the ICU to allow aerodynamic loads to drive the transcowls back to the fully deployed position.

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GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

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Page 43 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

THRUST REVERSER STOW STAGE Purpose (2.B.b) When the stow signal is sent from flight deck the transcowls are stowed. Operation (3.C.c) To stow the cowls, the FADEC commands the DCU solenoid to be deenergized and after a time delay, the ICU solenoid to be energized. A 10-second time delay in the SPDA maintains power to the cowl lock to allow for health checks of the cowl lock solenoid. The control valve spring in the DCU is sprung (and pressure assisted) into the stow direction allowing pressure to the actuators’ retract area and venting their heads to return. The internal actuator locks automatically re-engage as the cowls are retracted to the stowed position. The ICU solenoid is then deenergized 20 seconds after stow and the system returns to the deenergized state.

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Page 44 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

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Page 45 Aug 05

GE AIRCRAFT ENGINES

CF34-10E

TRAINING MANUAL

THRUST REVERSER SYSTEM INDICATION Identification (1.A.a) For all Thrust Reverser indications the icon REV will be used. On the EICAS Page (Main Page) the REV indication is located close to the N1/ITT dial for each engine. Purpose (2.B.b) The T/R indicating system supplies thrust reverser translating cowl position data to the pilot. Operation (3.C.c) There can be two conditions during a thrust reverser operation: - Normal operation - Abnormal operation Normal Operation With the thrust reverser in the stowed and locked position, but thrust reverser not locked out (not inhibited) there will be no indication. 4 4 0 1 5 0 0 0 0 0 8 7 D S

A thrust reverser in transition (stowing or deploying) is indicated by amber “REV”.

With the thrust reverser locked out (by the maintenance crew by flipping the inhibit lever switch on the ICU to the ‘On’ position and installing the two thrust reverser translating cowl locking pins) there will be a white “Ex REV INHIBIT” status message. Abnormal Operation A red "REV" icon, as well as the Ex REV DEPLOYED message are displayed on the EICAS when there is an uncommanded deployment in-flight or on the ground. Other abnormal operations which are indicated on the EICAS include: Abnormal Condition Uncommanded stow Failure to deploy Failure to stow Inadvertent deployment

Indication and Message Ex REV FAIL Ex REV FAIL Red REV icon & Ex REV DEPLOYED warning message Red REV icon & Ex REV DEPLOYED warning message

The red ENGX REV message indicates that the TR has faulted and the FADEC will command “return to idle”.

When fully deployed on a command by the pilot, the indication is a green “REV”.



Page 46 Sept 05

CF34-10E LM June 09 print.pdf

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