texto turbinas a gas 2013 001

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UMSS – FCYT

Maquinas Termicas II

Doc. Ing. Pedro Triveño H.

Issue 2 – April 2003

GAS TURBINE

ENGINES

CONTENTS

1 FUNDAMENTALS ........................................................................ 1-1

1.1 WORK, POWER & ENERGY ........................................................ 1-11.1.1 Work ................................................................................ 1-11.1.2 Power .............................................................................. 1-11.1.3 Energy ............................................................................. 1-2

1.2 FORCE AND MOTION .................................................................. 1-31.2.1 Force ............................................................................... 1-31.2.2 Velocity ............................................................................ 1-31.2.3 Acceleration .................................................................... 1-4

1.3 PRINCIPLES OF JET PROPULSION .......................................... 1-41.3.1 Thrust Calculation. .......................................................... 1-4

1.4 GAS TURBINES ............................................................................ 1-6

1.5 THE BRAYTON CYCLE................................................................ 1-7

1.6 CHANGES IN TEMPERATURE, PRESSURE AND VELOCITY . 1-91.6.1 Temperature and Pressure ............................................. 1-91.6.2 Velocity and Pressure ..................................................... 1-101.6.3 How The Changes are Obtained. .................................. 1-10

1.7 DUCTS AND NOZZLES................................................................ 1-10Continuity equation. ...................................................................... 1-10

1.7.2 Incompressible fluid flow. ................................................ 1-111.7.3 Bernoulli’s Theorem ........................................................ 1-111.7.4 Total energy. ................................................................... 1-12

1.8 CONTINUITY EQUATION AND BERNOULLI’S THEOREM ....... 1-131.8.1 Incompressible fluid. ....................................................... 1-131.8.2 Gas Laws ........................................................................ 1-15

1.9 SUBSONIC AIRFLOW THROUGH DIVERGENT AND CONVERGENT DUCTS 1-16

Divergent Duct ............................................................................... 1-161.9.2 Convergent Duct ............................................................. 1-16

SONIC AIRFLOW THOUGH DIVERGENT AND CONVERGENT DUCTS 1-17

1.11 THE WORKING CYCLE ON A PRESSURE VOLUME DIAGRAM 1-18

1.12 ENGINE CONFIGURATIONS. ...................................................... 1-191.12.1 Reaction engines ............................................................ 1-191.12.2 Power Engines ................................................................ 1-21

2 ENGINE PERFORMANCE ........................................................... 2-1

2.1 METHOD OF CALCULATING THE THRUST FORCES .............. 2-1

2.2 CALCULATING THE THRUST OF THE ENGINE ....................... 2-22.2.1 Comparison between thrust and horse-power ............... 2-6

2.3 ENGINE THRUST IN FLIGHT ...................................................... 2-72.3.1 Effect of forward speed ................................................... 2-92.3.2 Effect of afterburning on engine thrust ........................... 2-11



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2.3.3 Effect of altitude .............................................................. 2-112.3.4 Effect of temperature ...................................................... 2-13

2.4 PROPULSIVE EFFICIENCY ......................................................... 2-14

2.5 FUEL CONSUMPTION AND POWER TO WEIGHT RELATIONSHIP 2-15

2.6 SPECIFIC FUEL CONSUMPTION ............................................... 2-162.6.1 SPECIFIC FUEL CONSUMPTION – DEFINITION ........ 2-16

2.7 FLAT RATING ............................................................................... 2-16

2.8 PERFORMANCE RATINGS ......................................................... 2-16

3 INLET ............................................................................................ 3-1

3.1 INTRODUCTION ........................................................................... 3-13.2 RAM COMPRESSION .................................................................. 3-1

3.2.1 Importance of Ram Compression ................................... 3-1

3.3 TYPES OF AIR INTAKES ............................................................. 3-23.3.1 PITOT INTAKES ............................................................. 3-23.3.2 DIVIDED ENTRANCE DUCT ......................................... 3-3

3.4 IDEAL INTAKE CONDITIONS ...................................................... 3-4

3.5 INTAKE ANTI-ICING ..................................................................... 3-53.5.1 Engine Hot Air Anti-icing ................................................. 3-53.5.2 Engine Electrical Anti-icing ............................................. 3-73.5.3 Oil Anti-ice ....................................................................... 3-8

4 COMPRESSORS .......................................................................... 4-14.1 COMPRESSORS GENERAL ....................................................... 4-1

4.2 CENTRIFUGAL FLOW ................................................................. 4-14.2.1 Operation ......................................................................... 4-3

4.3 THE AXIAL FLOW COMPRESSOR ............................................. 4-5Operation ....................................................................................... 4-6

4.4 COMPRESSOR STALL AND SURGE ......................................... 4-134.4.1 Airflow Control System Principles ................................... 4-134.4.2 Compressor Characteristics ........................................... 4-174.4.3 Effect of Temperature on the Operating Point of the Airflow Control System 4-18

4.5 AIR FLOW CONTROL SYSTEM – OPERATION ........................ 4-20

4.6 AEROFOIL THEORY AND THE AXIAL FLOW COMPRESSOR (CONTINUED) 4-25

4.6.1 Speed of Airflow Over Blades ......................................... 4-254.6.2 Angle of Attack ................................................................ 4-25Some Important Points about Angle of Attack .............................. 4-26

4.7 APPLICATION TO THE AXIAL FLOW COMPRESSOR .............. 4-274.7.1 Compressor RPM ............................................................ 4-274.7.2 Common Causes of Compressor Stall ........................... 4-274.7.3 Stagger Angle and End Bend ......................................... 4-274.7.4 Recent innovations ......................................................... 4-27



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4.8 AIRFLOW CONTROL ................................................................... 4-29

4.9 AIR BLEED VALVES (SUMMARY) .............................................. 4-29

4.10 VARIABLE INTAKE GUIDE VANES (SUMMARY) ...................... 4-29

4.11 MULTI-SPOOL COMPRESSORS (SUMMARY) .......................... 4-29

4.12 COMPARING THE FEATURES OF CENTRIFUGAL AND AXIAL FLOWCOMPRESSORS ...................................................................................... 4-30

4.12.1 Centrifugal ....................................................................... 4-304.12.2 Axial Flow ........................................................................ 4-30

5 COMBUSTION SECTION ............................................................. 5-1

5.1 INTRODUCTION ........................................................................... 5-1

5.2 COMBUSTION PROCESS ........................................................... 5-1

5.3 FUEL SUPPLY .............................................................................. 5-3

5.4 TYPES OF COMBUSTION CHAMBER ....................................... 5-45.4.1 Multiple combustion chamber ......................................... 5-45.4.2 Tubo-annular combustion chamber ................................ 5-6(Also known as Can-annular or Cannular.) .................................. 5-65.4.3 Annular combustion chamber ......................................... 5-75.4.4 Reverse Flow Combustion Chamber .............................. 5-9

5.5 COMBUSTION CHAMBER PERFORMANCE ............................. 5-105.5.1 Combustion intensity ....................................................... 5-10

5.6 COMBUSTION EFFICIENCY ....................................................... 5-11

5.7 COMBUSTION STABILITY ........................................................... 5-11

5.8 POLLUTION CONTROL ............................................................... 5-125.8.1 Introduction...................................................................... 5-125.8.2 Sources of Pollution ........................................................ 5-12

5.9 EMISSIONS .................................................................................. 5-12

5.10 MATERIALS .................................................................................. 5-14

6 TURBINE SECTION ..................................................................... 6-1

6.1 INTRODUCTION ........................................................................... 6-1

6.2 ENERGY TRANSFER FROM GAS FLOW TO TURBINE ........... 6-5

6.3 CONSTRUCTION ......................................................................... 6-86.3.1 Nozzle guide vanes ......................................................... 6-86.3.2 Turbine discs ................................................................... 6-96.3.3 Turbine blades ................................................................ 6-96.3.4 Dual alloy discs ............................................................... 6-11

6.4 COMPRESSOR-TURBINE MATCHING ...................................... 6-11

6.5 MATERIALS .................................................................................. 6-116.5.1 Nozzle guide vanes ......................................................... 6-116.5.2 Turbine discs ................................................................... 6-116.5.3 Turbine blades ................................................................ 6-12

6.6 DYNAMIC BALANCING PRINCIPLES ......................................... 6-16



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6.6.1 Introduction...................................................................... 6-166.6.2 Centrifugal Force ............................................................. 6-176.6.3 Causes of Unbalance ...................................................... 6-186.6.4 Objective of Balancing .................................................... 6-206.6.5 Definition of Unbalance ................................................... 6-206.6.6 Fan Balancing ................................................................. 6-23

7 EXHAUST ..................................................................................... 7-1

7.1 INTRODUCTION ........................................................................... 7-1

EXHAUST GAS FLOW ............................................................................. 7-3

7.3 CONSTRUCTION AND MATERIALS ........................................... 7-7

7.4 NOISE REDUCTION ..................................................................... 7-87.4.1 Sources of Engine Noise ................................................ 7-8

7.5 THRUST REVERSAL ................................................................... 7-187.5.1 Introduction...................................................................... 7-187.5.2 Requirement for Thrust Reversal ................................... 7-187.5.3 Layout and Operation of Typical Thrust Reversing Systems 7-197.5.4 Safety Features ............................................................... 7-22CFM 56 Thrust Reverser for Boeing 737-300 .............................. 7-22

8 BEARINGS, SEALS AND GEARBOXES ..................................... 8-1

8.1 BEARINGS .................................................................................... 8-18.1.1 Introduction...................................................................... 8-1

8.1.2 Ball Bearings ................................................................... 8-18.1.3 Roller Bearings ................................................................ 8-18.1.4 Other types of bearings ................................................... 8-1

8.2 BEARING CHAMBER OR SUMP ................................................. 8-38.2.1 Lubrication ....................................................................... 8-38.2.2 Sealing ............................................................................ 8-38.2.3 Thread Seals ................................................................... 8-48.2.4 Carbon Seal .................................................................... 8-58.2.5 Spring Ring Seal ............................................................. 8-58.2.6 Hydraulic Seal ................................................................. 8-6

8.3 ACCESSORY DRIVE GEARBOXES ............................................ 8-78.3.1 Introduction...................................................................... 8-78.3.2 Internal gearbox .............................................................. 8-78.3.3 Radial driveshaft ............................................................. 8-108.3.4 Direct drive ...................................................................... 8-108.3.5 Gear train drive ............................................................... 8-108.3.6 Intermediate gearbox ...................................................... 8-108.3.7 External gearbox ............................................................. 8-118.3.8 Auxiliary gearbox ............................................................. 8-128.3.9 Construction and Materials ............................................. 8-14

9 LUBRICANTS AND FUEL ............................................................ 9-1

9.1 GAS TURBINE FUEL PROPERTIES AND SPECIFICATION ..... 9-1



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9.2 FRACTIONAL DISTILLATION ...................................................... 9-1

9.3 PROPERTIES ............................................................................... 9-39.3.1 Ease of Flow ................................................................... 9-39.3.2 Ease of Starting ............................................................... 9-39.3.3 Complete Combustion .................................................... 9-39.3.4 Calorific Value ................................................................. 9-49.3.5 Corrosive Properties ....................................................... 9-49.3.6 Effects of By-Products of Combustion ............................ 9-59.3.7 Fire Hazards .................................................................... 9-59.3.8 Vapour Pressure ............................................................. 9-69.3.9 Fuel Boiling and Evaporation Losses ............................. 9-6

9.3.10 Methods of Reducing or Eliminating Fuel Losses .......... 9-69.3.11 Fuel additives .................................................................. 9-89.3.12 Safety precautions .......................................................... 9-8

9.4 GAS TURBINE OIL PROPERTIES AND SPECIFICATIONS ...... 9-99.4.1 Viscosity .......................................................................... 9-99.4.2 Hydro-Dynamics or Fluid Film Lubrication ..................... 9-99.4.3 Boundary Lubrication ...................................................... 9-10

9.5 LUBRICATING OILS ..................................................................... 9-10

9.6 TURBINE OILS ............................................................................. 9-119.6.1 First Generation Synthetic Oils ....................................... 9-129.6.2 Second Generation Synthetic Oils .................................. 9-129.6.3 Third Generation Synthetic Oils ...................................... 9-129.6.4 Safety Precautions .......................................................... 9-13

10 LUBRICATION SYSTEMS ........................................................... 10-1

10.1 INTRODUCTION ........................................................................... 10-1

10.2 BEARINGS .................................................................................... 10-1

10.3 ENGINE LUBRICATION SYSTEMS ............................................. 10-510.3.1 Pressure Relief Valve Re-circulatory System ................. 10-510.3.2 Recirculatory Oil System – Full Flow Type ..................... 10-810.3.3 Advantages of Full Flow Lubrication ............................... 10-810.3.4 Expendable System ........................................................ 10-10

10.4 MAIN COMPONENTS .................................................................. 10-11

10.4.1 Oil Tank ........................................................................... 10-1110.4.2 Oil Pumps ........................................................................ 10-1210.4.3 oil cooling ........................................................................ 10-1410.4.4 Pressure Filter ................................................................. 10-1510.4.5 Last Chance Filter ........................................................... 10-1710.4.6 Scavenge Oil Strainers ................................................... 10-1710.4.7 Magnetic Chip Detector .................................................. 10-1810.4.8 De-aerator ....................................................................... 10-1810.4.9 Centrifugal Breather ........................................................ 10-19Pressure Relief Valve .................................................................... 10-1910.4.11 By-Pass Valve ................................................................. 10-20



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10.5 INDICATIONS AND WARNINGS ................................................. 10-2110.5.1 Low Pressure Warning Lamp ......................................... 10-2110.5.2 Oil Pressure, temperature and quantity indication ......... 10-21

10.6 OIL SEALS .................................................................................... 10-21

10.7 SERVICING ................................................................................... 10-21

11 ENGINE FUEL CONTROL SYSTEMS ......................................... 11-1

11.1 INTRODUCTION ........................................................................... 11-1

11.2 PURPOSE OF THE ENGINE FUEL SYSTEM ............................. 11-1

11.3 LAYOUT OF TYPICAL SYSTEM COMPONENTS ...................... 11-311.3.1 Aircraft Mounted Components ........................................ 11-311.3.2 The Engine LP fuel system ............................................. 11-311.3.3 The Engine HP Fuel System .......................................... 11-3

11.4 FACTORS GOVERNING FUEL REQUIREMENTS ..................... 11-5

11.5 REQUIREMENTS OF THE ENGINE FUEL SYSTEM ................. 11-5

11.6 ENGINE FUEL SYSTEM COMPONENTS ................................... 11-5

11.7 FUEL PUMPS ............................................................................... 11-511.7.1 Fuel Pump Requirements ............................................... 11-511.7.2 Plunger-type Fuel Pump ................................................. 11-611.7.3 Gear-Type Fuel Pump .................................................... 11-7

11.8 FUEL FLOW CONTROL ............................................................... 11-711.8.1 Basic Flow Control System ............................................. 11-8

11.9 HYDRO-MECHANICAL CONTROL UNITS ................................. 11-1011.9.2 Barometric Controls ........................................................ 11-1111.9.3 Proportional Flow Control. .............................................. 11-1311.9.4 Acceleration Control Units .............................................. 11-14

11.10 ENGINE PROTECTION DEVICES ............................................... 11-1811.10.1 Top Temperature Limiter. ............................................... 11-1811.10.2 Power Limiter. ................................................................. 11-1811.10.3 Overspeed Governor. ..................................................... 11-19

BURNERS ................................................................................................ 11-2111.11.1 Atomiser Burners ............................................................ 11-2111.11.2 Vaporising Burners ......................................................... 11-26

11.11.3 Combustion and Airflow .................................................. 11-2811.12 ELECTRONIC ENGINE CONTROL SYSTEMS ........................... 11-30

11.12.1 Supervisory Electronic Engine Control ........................... 11-3011.12.2 FUEL CONTROL ............................................................ 11-3211.12.3 General ............................................................................ 11-3211.12.4 Full-Authority Digital Electronic Control (FADEC) .......... 11-36

12 AIR SYSTEMS .............................................................................. 12-1

12.1 INTRODUCTION ........................................................................... 12-1

12.2 INTERNAL COOLING AIRFLOW ................................................. 12-212.2.1 Low Pressure Air ............................................................. 12-2



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12.2.2 Intermediate Pressure Air ............................................... 12-212.2.3 High Pressure Air ............................................................ 12-212.2.4 Differential Pressure Seals ............................................. 12-3

12.3 SEALING ....................................................................................... 12-3

12.4 COOLING. ..................................................................................... 12-5

12.5 TURBINE CASE COOLING – DESCRIPTION AND OPERATION 12-912.5.1 Passive Clearance Control System. Figure 12.7. .......... 12-912.5.2 Active Clearance Control System. Figure 12.8. ............. 12-1012.5.3 Low Pressure Turbine Clearance Control Valve ............ 12-11

12.6 EXTERNAL COOLING .................................................................. 12-13

12.6.1 External skin of aero-engine. .......................................... 12-1312.6.2 Cooling of Accessories ................................................... 12-14

12.7 HP AIR FOR AIRCRAFT SERVICES. .......................................... 12-15External Air Tappings .................................................................... 12-15

12.8 ANTI-ICING SYSTEMS................................................................. 12-18

13 STARTING AND IGNITION SYSTEMS ........................................ 13-1

13.1 BASIC PRINCIPLES OF GAS TURBINE ENGINE STARTING SYSTEMS 13-113.1.1 Purpose ........................................................................... 13-113.1.2 Essential Starting Requirements .................................... 13-1

STARTER MOTORS ................................................................................ 13-213.2.1 Electrical Starter Motor ................................................... 13-3

13.2.2 Electric Starter/Generator ............................................... 13-313.2.3 Safety Interlocks .............................................................. 13-413.2.4 Air Turbo Starters ............................................................ 13-5

13.3 A300 STARTING SYSTEM ........................................................... 13-813.3.1 GE 6-50 Starting Procedure ........................................... 13-8

13.4 IGNITION SYSTEMS .................................................................... 13-1213.4.1 High Energy Ignition Unit ................................................ 13-1213.4.2 Igniter Plug ...................................................................... 13-1413.4.3 Servicing the Ignition System ......................................... 13-14

14 ENGINE INDICATION SYSTEMS ................................................ 14-1

14.1 INTRODUCTION. .......................................................................... 14-1

14.2 ENGINE SPEED INDICATORS. ................................................... 14-3

14.3 THRUST INDICATION .................................................................. 14-714.3.1 Engine Pressure Ratio.EPR. .......................................... 14-714.3.2 Torque indication ............................................................. 14-914.3.3 Phase comparison Torquemeter .................................... 14-12

14.4 EXHAUST GAS TEMPERATURE ................................................ 14-1314.4.1 Thermocouples ............................................................... 14-13

14.5 FUEL FLOW METERING ............................................................. 14-17

14.6 OIL ................................................................................................. 14-2014.6.1 The Oil Pressure Indicator .............................................. 14-20



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14.6.2 Oil pressure warning light ............................................... 14-21Oil Temperature. ........................................................................... 14-2214.6.4 Oil Quantity...................................................................... 14-23

14.7 VIBRATION ................................................................................... 14-24

14.8 WARNING LIGHTS ....................................................................... 14-24

15 THRUST AUGMENTATION ......................................................... 15-1

15.1 INTRODUCTION ........................................................................... 15-1

15.2 WATER INJECTION ..................................................................... 15-115.2.1 Effects on Engine Power................................................. 15-115.2.2 Methods of Applying Water/Methanol ............................. 15-1

15.2.3 Compressor Intake Injection (Turbo Prop) ..................... 15-215.2.4 Combustion Chamber Injection System ......................... 15-4

15.3 RE-HEAT (AFTER BURNING) ..................................................... 15-615.3.1 Purpose ........................................................................... 15-615.3.2 Revision of Thrust ........................................................... 15-615.3.3 Re-heat and By-pass Engines ........................................ 15-615.3.4 The Advantage of Re-heat .............................................. 15-615.3.5 The disadvantages of Re-heat ........................................ 15-715.3.6 Propelling Nozzles .......................................................... 15-715.3.7 Re-heat Nozzles .............................................................. 15-815.3.8 The Re-heat Jet Pipe ...................................................... 15-10

16 TURBO-PROP ENGINES ............................................................. 16-116.1 INTRODUCTION ........................................................................... 16-1

16.2 TYPES OF TURBO-PROP ENGINES .......................................... 16-116.2.1 Coupled Power Turbine .................................................. 16-116.2.2 Free Power Turbine ........................................................ 16-216.2.3 Compounded Engine ...................................................... 16-3

16.3 REDUCTION GEARING ............................................................... 16-316.3.1 Simple Spur ‘Epicyclic’ .................................................... 16-416.3.2 Compound Spur Epicyclic ............................................... 16-616.3.3 Gear Train/Epicyclic ........................................................ 16-7

16.4 TURBO-PROP PERFORMANCE ................................................. 16-7

16.5 TURBO-PROP ENGINE CONTROL ............................................ 16-716.5.1 Integrated Control of RPM and Fuel Flow ...................... 16-816.5.2 Direct Control of Fuel Flow ............................................. 16-816.5.3 Direct Control of Blade Angle (Beta Control) .................. 16-8

16.6 ENGINE AND PROPELLER CONTROLS .................................... 16-9

16.7 CONTROL OUTSIDE NORMAL FLIGHT RANGE ....................... 16-9

16.8 PROPELLER CONTROL .............................................................. 16-916.8.1 Constant Speed Unit ....................................................... 16-1016.8.2 Manual and Automatic Feathering Controls ................... 16-1016.8.3 Fixed and Removable Stops ........................................... 16-15

16.9 OVERSPEED SAFETY DEVICES ................................................ 16-16



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17 TURBOSHAFT ENGINES ............................................................ 17-1

17.1 INTRODUCTION. .......................................................................... 17-1

17.2 FUEL CONTROL SYSTEM .......................................................... 17-4

17.3 ARRANGEMENTS ........................................................................ 17-6

17.4 DRIVE SYSTEMS ......................................................................... 17-10

17.5 COUPLINGS ................................................................................. 17-13

18 AUXILLIARY POWER UNITS ...................................................... 18-1

18.1 INTRODUCTION ........................................................................... 18-1

18.2 GENERAL ARRANGEMENTS AND CONFIGURATION ............. 18-3

18.2.1 Inlet Duct Arrangement ................................................... 18-718.2.2 Exhaust Duct Arrangement ............................................. 18-9

18.3 THE APU ENGINE ........................................................................ 18-10

18.4 FUEL CONTROL........................................................................... 18-12Mechanical Fuel Control ............................................................... 18-1218.4.2 Speed Control ................................................................. 18-1818.4.3 Mechanical Fuel Control Unit Operation ........................ 18-1918.4.4 Electronic APU Fuel Control ........................................... 18-2018.4.5 Electro/mechanical Fuel Control (FIGURE 18.26) ......... 18-21

18.5 APU OIL SYSTEM ........................................................................ 18-23

18.6 APU BLEED AIR SYSTEMS ......................................................... 18-25

18.6.1 direct from engine compressor ....................................... 18-2518.6.2 SEPARATE LOAD COMPRESSOR ............................... 18-27

18.7 BAY COOLING .............................................................................. 18-2818.7.1 Ram Air Cooling .............................................................. 18-2818.7.2 Fan Air Cooling ............................................................... 18-28

18.8 APU POWERPLANT INSTALLATION.......................................... 18-32

18.9 APU STARTING SEQUENCE ...................................................... 18-34

19 POWERPLANT INSTALLATION ................................................. 19-1

19.1 NACELLES OR PODS .................................................................. 19-119.1.1 Cowlings .......................................................................... 19-119.1.2 Firewalls .......................................................................... 19-4

19.1.3 Cooling ............................................................................ 19-619.1.4 Acoustic Linings .............................................................. 19-819.1.5 Abradable Linings ........................................................... 19-11

19.2 ENGINE MOUNTS ........................................................................ 19-1219.2.1 Wing Pylon Mounted Engine (Turbofan) ........................ 19-1219.2.2 Wing Mounted Engine (Turboprop) ................................ 19-1419.2.3 Rear Fuselage Engine Turbofan.(Figure 19.14/15.) ...... 19-16

19.3 ENGINE DRAINS. ......................................................................... 19-1819.3.1 Controlled Drains ............................................................ 19-1819.3.2 Uncontrolled Drains ........................................................ 19-20

19.4 ENGINE CONTROLS ................................................................... 19-22



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19.4.1 Throttle Control Mechanical ............................................ 19-2219.4.2 Turbofan Engine Controls. .............................................. 19-2219.4.3 Turboprop Engine Controls ............................................. 19-24

19.5 ENGINE BUILD UNIT ................................................................... 19-2919.5.1 Turbofan Engine .............................................................. 19-29

19.6 FIRE PREVENTION – BAYS OR ZONES .................................... 19-38

19.7 INSTALLING AND REMOVING ENGINES. ................................. 19-4019.7.1 Removal .......................................................................... 19-4019.7.2 Fitting ............................................................................... 19-4819.7.3 Turbo Prop Engine Removal/Fit. .................................... 19-4819.7.4 Flight Transit ................................................................... 19-48

20 FIRE PROTECTION SYSTEMS ................................................... 20-1

20.1 FIRE DETECTORS ....................................................................... 20-1

20.2 FIRE WIRE SYSTEMS ................................................................. 20-320.2.1 Resistance Type ............................................................. 20-320.2.2 Capacitance Type ........................................................... 20-320.2.3 Gas Operation Fire Wire ................................................. 20-420.2.4 Single Loop ..................................................................... 20-520.2.5 Dual Loop ........................................................................ 20-5Dual Loop Systems ....................................................................... 20-6

20.3 FIRE AND LOOP FAULT INDICATION (E.C.A.M.) ...................... 20-8

20.4 FIRE SUPPRESSION ................................................................... 20-920.4.1 Types of Fire Suppression System ................................. 20-11One Shot System .......................................................................... 20-1120.4.2 Two Shot System (shared extinguishers) ....................... 20-1220.4.3 Two Shot System (Single Head extinguishers) .............. 20-14

20.5 EXTINGUISHERS ......................................................................... 20-1620.5.1 Operating Head ............................................................... 20-1720.5.2 Safety Discharge ............................................................. 20-1720.5.3 Discharge Tube Configuration ........................................ 20-1820.5.4 Operating Time ............................................................... 20-1920.5.5 Extinguishant ................................................................... 20-19

20.6 INDICATIONS OF FIRE DETECTION.......................................... 20-20

20.6.1 Fire T Handle .................................................................. 20-2020.6.2 Fire Bell ........................................................................... 20-2020.6.3 Fire Detection Test .......................................................... 20-22

20.7 DISCHARGE INDICATORS ......................................................... 20-2320.7.1 Mechanical Indicators ..................................................... 20-2320.7.2 Electrical Indicators ......................................................... 20-23

20.8 CARTRIDGES OR SQUIBS ......................................................... 20-2420.8.1 Life Control of Squibs ...................................................... 20-24

INTENTIONALLY BLANK ......................................................................... 20-26

21 ENGINE MONITORING AND GROUND OPERATIONS. ............. 21-1



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21.1 PROCEDURES FOR STARTING AND GROUND RUNNING. ... 21-1

21.2 STARTING .................................................................................... 21-3

21.3 UNSATISFACTORY STARTS ...................................................... 21-7

21.4 ENGINE STOPPING. .................................................................... 21-8

21.5 ENGINE FIRES ............................................................................. 21-9

21.6 INTERPRETATION OF ENGINE POWER OUTPUTS AND PARAMETERS. 21-10

21.7 TREND MONITORING. ................................................................ 21-2221.7.1 On Ground Monitoring .................................................... 21-2421.7.2 Air Washed Components ................................................ 21-2421.7.3 Oil Washed Components ................................................ 21-32

21.7.4 Inspections ...................................................................... 21-36

22 ENGINE STORAGE AND PRESERVATION. ............................... 22-1

22.1 STORAGE AND TRANSIT ........................................................... 22-122.1.1 Fuel System Inhibiting. ................................................... 22-122.1.2 Packing. ........................................................................... 22-222.1.3 Storage. ........................................................................... 22-3



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1 FUNDAMENTALS

1.1 WORK, POWER & ENERGY

Work, power, and energy are all interrelated. Work is the amount of movement agiven force causes; energy is the ability to do work, and power is the rate of doingwork.

1.1.1 WORK

In its technical sense, work is the product of force and distance, and work is doneonly when a force causes movement. We can see this by the formula:

Work = Force x Distance

We normally measure distance in feet or inches, and force in pounds or ounces.

This allows us to measure work in foot-pounds or inch-ounces.

Example:

To find the amount of work done when a 500 pound load is lifted for a distance of6 feet, we can use the formula:

Work = Force x Distance

= 500 X 6

= 3,000 foot-pounds

1.1.2 POWER

The rate of doing work is called power, and it is defined as the work done in unittime. As a formula, this would be:

power = work donetime taken

Power is expressed in several different units, such as the watt, ergs per second,and foot-pounds per second. The most common unit of power in general use inthe United States is the horsepower. One horsepower (hp) is equal to 550 ft-lb’sor 33000 ft-1b/min. In the metric system the unit of power is the watt (W) or thekilowatt (kW). One hp is equal to 746 watts; and 1 kW = 1.34 hp.

Example:

To compute the power necessary to raise an elevator containing 10 persons adistance of 100 ft in 5 s (assuming the loaded elevator weighs 2500 lb), proceedas follows:

Power = work done = 2500 x 100 = 50,000 ft-lb’s/secTime taken 5

Since 1hp = 550 ft-lb’s/sec then required hp = 50,000550

= 90.9 hp (67.81 kw assuming no friction losses)



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1.1.3 ENERGY

The term energy may be defined as the capacity for doing work. There are twoforms of energy: potential energy and kinetic energy.

1.1.3.1 POTENTIAL ENERGY

Potential energy is the stored energy possessed by a system, because of therelative positions of the components of that system. If work done raises an objectto a certain height, energy will be stored in that object in the form of thegravitational force. This energy, waiting to be released is called potential energy.The amount of potential energy a system possesses is equal to the work done onthe system previously.

Potential energy can be found in forms other than weights and height. Electrically

charged components contain potential (electrical) energy because of their positionwithin an electric field. An explosive substance has chemical potential energy thatis released in the form of light, heat and kinetic energy, when detonated.

Example :

A weight of 50 pounds is raised 5 feet. Using the formula:

Potential Energy = Force x Distance

= 50 x 5

= 250 ft-lb’s.

Note: That energy is expressed in the same units as those used for work and in allcases energy is the product of force x distance.

1.1.3.2 KINETIC ENERGY

Kinetic energy is the energy possessed by an object, resulting from the motion ofthat object. The magnitude of that energy depends on both the mass and speedof the object. This is demonstrated by the simple equation:

Energy =½mv2 or w v2 2g

where m = mass, v = velocity (in feet or metres per second), w = weight, g =

gravity (32 ft/sec2

or 9.81m/sec2

). All forms of energy convert into other forms by appropriate processes. In thisprocess of transformation, either form of energy can be lost or gained but the totalenergy must remain the same.

Example:

A weight of 50lbs dropped from a height of 5 ft has kinetic energy of

KE = 50 x 252 x 32

= 19.53 ft-lb’s



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1.2 FORCE AND MOTION

1.2.1 FORCEForce may be defined as a push or a pull upon an object. In the English systemthe pound (1b) is used to express the value of a force. For example, we say that aforce of 30 lb is acting upon a hydraulic piston.

A unit of force in the metric system is the newton (N). The newton is the forcerequired to accelerate a mass of 1 kilogram (kg) 1 meter per second per second(m/s2).

The dyne (dyn) is also employed in the metric system as a unit of force. One dyneis the force required to accelerate a mass of 1g 1 centimetre per second persecond (cm/s2). One newton is equal to 100,000 dynes (0.225 Ib).

1.2.2 VELOCITY

It is common to find people confusing the terms velocity and speed whendescribing how fast an object is moving. The difference is that speed is a scalarquantity, whilst the term velocity refers to both speed and direction of an object.The full definition of velocity is that it is the rate at which its position changes, overtime, and the direction of the change.

The simple diagram below shows how an aircraft, which flies the irregular pathfrom 'A' to 'B' in an hour, (a speed of 350 mph), has an actual velocity of 200 mphin an East-Northeast direction.

A

B

C

N

200 Ml (322 Km)

350 Ml (563 Km)

Path of Aircraft

Diagram Showing Difference Between Velocity and Speed

Figure 1.1.



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1.2.3 ACCELERATION

This term describes the rate at which velocity changes. If an object increases inspeed, it has positive acceleration; if it decreases in speed, it has negativeacceleration. A reference to Newton's Second law of Motion will explain theprinciples of acceleration. Acceleration can be in a straight line, which is referredto a linear acceleration and it can apply to rotating objects whose speed of rotationis increasing, (or decreasing), when it is called angular acceleration.

1.3 PRINCIPLES OF JET PROPULSION

Newton’s Laws of Motion. To understand the basic principles of jet propulsion it isnecessary to understand the practical application of Sir Isaac Newton's Laws ofMotion. There are three laws.

1. The First Law States. A mass will remain stationary until acted upon by aforce. If the mass is already moving at a constant speed in a straight line, it will.continue to move at that constant speed in a straight line until acted upon by aforce.

2. The Second Law States. When a force acts on a mass the mass willaccelerate in the direction in which the force acts.

3. The Third Law States. To every action there is an equal and opposite reaction.

The function of any propeller or gas turbine engine is to produce THRUST, (or apropulsion force), by accelerating a mass of air or gas rearwards. If we apply

Newton's Laws of Motion to aircraft propulsion it can be said that:- a FORCE must be applied in order to accelerate the mass of air or gas: first

law,

the acceleration of the mass is proportional to the force applied: second law,

there must be an equal and opposite reaction, in our case this is THRUST, aforward acting force: third law.

1.3.1 THRUST CALCULATION.

The amount of thrust produced depends upon two things:-

the MASS of air which is moved rearwards in a given time, the ACCELERATION imparted to the air.

It can be expressed as:- Thrust = Mass x Acceleration

The MASS is defined as “the quantity of matter in a body".

It is expressed as W

gWhere:- W = the weight of the body (in lb’s or newtons) and

g = the gravitational constant (taken as 32 ft/sec/sec or 9.81 m/sec2)

The ACCELERATION imparted to the air is the difference between its inlet andoutlet velocity.



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If we let: -V2 = the air velocity at exit (in ft/sec/sec or 9.81m/sec2)

and

V1 = the air velocity at inlet (in ft/sec/sec or 9.81m/sec2)

It may be expressed as V2 – V1

Taking these expressions for Mass and Acceleration, the thrust produced by anengine or propeller can be calculated from the following formula:-

THRUST = 12 V-Vg

W

Example 1.The airflow through a propeller is 256 lbs/sec, Inlet velocity 0 ft/sec, outlet velocity700 ft/sec.

Thrust developed will be:

THRUST = 12 V-Vg

W

THRUST = 256 x (700 – 0)32

= 5600 lbs

Example 2.

The mass airflow through a gas turbine engine is 128lbs/sec, inlet velocity is 0ft/sec, outlet velocity is 1400 ft/sec. Using the formula :

THRUST = 128 x (1400 – 0)

32

= 5600lbs

By comparing both examples, you can see that the gas turbine produced the samethrust as the propeller by giving a greater acceleration to a smaller mass. It canbe said that a propeller accelerates a large mass slowly whilst the gas turbineproduces the same thrust by giving a greater acceleration to a smaller mass.

Note that in both of the examples the inlet velocity was zero ft/sec. The aircraftwas stationary so the thrust produced is referred to as STATIC THRUST.



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1.4 GAS TURBINES

A gas turbine engine is essentially a heat engine using a mass of air as a workingfluid to provide thrust. To achieve this, the mass of air passing through the enginehas to be accelerated, which means that the velocity, (or kinetic energy), of the airis increased. To obtain this increase, the pressure energy is first of all increased,followed by the addition of heat energy, before final conversion back to kinetic energy in the form of a high velocity jet efflux.

The simplest form of gas turbine engine is the turbojet engine, which has threemajor parts; the compressor, the combustion section and the turbine. A shaftconnects the compressor and the turbine to form a single, rotating unit. Theseengines produce thrust in the manner described in the Brayton Cycle.

The simplest turbojet engine is the unit shown below with a singlecentrifugal(Double Entry)compressor and a single stage turbine. This type ofengine can still be found in certain special installations but generally, they havebeen superseded by engines with axial compressors and multiple stage turbines.The advantages and disadvantages of the two types of compressor will bediscussed in depth later in this module

Simple Centrifugal Gas Turbine (Derwent)Figure 1.2.



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1.5 THE BRAYTON CYCLE

The working cycle of the gas turbine engine is similar to that of the four-strokepiston engine. There is induction, compression, ignition and exhaust in bothcases, although the process is continuous in a gas turbine. Also, the combustionin a piston engine occurs at a constant volume, whilst in a gas turbine engine itoccurs at a constant pressure.

The cycle, upon which the gas

turbine engine functions, in itssimplest form, is the Braytoncycle, which is represented bythe pressure/volume diagram,shown in figure 1.4.

The Working Cycle.Figure 1.3.

The Brayton Cycle.Figure 1.4.



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• The air entering the engine is compressed.• Heat is added to the air by burning fuel at a constant pressure, thereby

considerably increasing the volume of the resulting gas.

• The gases resulting from combustion expand through the turbine, whichconverts some of the energy in the expanding gases into mechanical energyto drive the compressor.

• The remainder of the expanding gases are propelled through the turbine and jet pipe back to the atmosphere where they provide the propulsive jet.

There are three main stages in the engine working cycle during which the changes

discussed occur:• During compression. Work is done on the air. This increases the pressure

and temperature and decreases the volume of air.

• During combustion. Fuel is added to the air and then burnt. This increasesthe temperature and volume of the gas, whilst the pressure remains almostconstant (the latter being arranged by design in a gas turbine engine).

• During expansion. Energy is taken from the gas stream to drive thecompressor via the turbine; this decreases the temperature and pressure,whilst the volume increases. The rapidly expanding gases are propelledthrough the turbine and jet pipe to give a final momentum that is much greater

than the initial momentum; it is this change in momentum which produces thepropulsive jet.



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1.6 CHANGES IN TEMPERATURE, PRESSURE AND VELOCITY .1.6.1 TEMPERATURE AND PRESSURE

The changes in temperature and pressure of the gases through a gas turbineengine are illustrated in Figure 1.5 The efficiency with which these changes aremade will determine to what extent the desired relations between pressure,temperature and velocity are obtained. The more efficient the compressor, thehigher is the pressure generated for a given work input - i.e. for a giventemperature rise of the gas. Conversely, the more efficiently the turbine uses theexpanding gas, the greater is the output of work for a given temperature drop ingas.

Gas Flow Through an Engine

Figure 1.5



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1.6.2 VELOCITY AND PRESSUREDuring the passage of the air (gas) through the engine, aerodynamic and energyrequirements demand changes in its velocity and pressure. For example, duringcompression a rise in the pressure of the air is required with no increase in itsvelocity. After the air has been heated and its internal energy increased bycombustion, an increase in the velocity of the gases is necessary to cause theturbine to rotate. Also at the propelling nozzle, a high velocity is required, for it isthe change in momentum of the air that provides the thrust on the aircraft. Localdecelerations of gas flow are also required - for example, in the combustionchambers to provide a low velocity zone for the flame.

1.6.3 HOW THE CHANGES ARE OBTAINED.The various changes in temperature, pressure and velocity are effected by meansof the ducts through which the air (gas) passes on its way through the engine.When a conversion from kinetic energy to pressure energy is required, the ductsare divergent in shape. Conversely, when it is required to convert the energystored in the combustion gases to velocity, a convergent nozzle is used. Thedesign of the passages and nozzles is of great importance, for upon their gooddesign depends the efficiency with which the energy changes are effected. Anyinterference with the smooth flow of gases creates a loss in efficiency and couldresult in component failure because of vibration caused by eddies or turbulence ofthe gas flow.

1.7 DUCTS AND NOZZLES

1.7.1 CONTINUITY EQUATION.

If we consider the machine to be an open-ended duct (Fig 1.6.), we find that themass flow per second will depend on the density of the fluid and the volumeflowing per sec:

Now volume flow = Area of duct x distance travelled (L)

Time (sec)

Open Ended Duct to Illustrate Continuity Equation

Figure 1.6.



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But the distance travelled per second = Velocity.

Therefore, Mass flow = density x area x velocity.This is known as the ‘continuity equation’ and it is true for any steady flow systemregardless of changes in the cross-sectional area of the duct.

1.7.2 INCOMPRESSIBLE FLUID FLOW.

Now consider an incompressible fluid as it flows through the duct system shown inthe fig. 1.7. We know that the mass flow is of a constant value and, naturally, asthe fluid enters the larger cross sectional area it will take up the new shape andthe initial volume will now occupy less length in the duct. Therefore, in a giventime, less distance is travelled and the velocity is reduced.

Thus we conclude that if the mass flow is to remain constant, as it must, anincrease in duct area must be accompanied by a reduction in flow velocity, and adecrease in duct area must bring about an increase in velocity; we can expressthis action as – velocity varies inversely with changes in duct area.

1.7.3 BERNOULLI’S THEOREM

This theorem can be related to the relationship between pressure and velocityexisting in the air flowing through a duct, such as a jet engine. The theorem states

that the total energy per unit mass is constant for a fluid moving inside a duct andthat total energy consists mainly of pressure energy and kinetic energy:

Pressure energy.

In gas or fluid flow the pressure energy is more often called ‘static pressure’ and itcan be defined as the pressure that would be felt by a body which was submergedin the medium (gas or fluid) and moving at the same velocity as the medium.

Duct SystemFigure 1.7.



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Kinetic energy.This kind of energy is more often called ‘dynamic pressure’ and this term is usedto define the extra pressure created by the movement of the medium. Dynamicpressure is proportional to ½ mass x velocity 2 (ie. ½mv2).

When the medium (gas or fluid) is moving, the total energy = static pressure +dynamic pressure.

Consider a duct which is filled with an incompressible fluid and pressurised fromone end by an external force (Fig 1.8.). The other end of the duct is sealed by avalve, which can be opened or closed, and a pressure gauge is fitted into the wallof the duct to indicate the static pressure (PS). With the valve closed, staticpressure and total energy are the same. However, when the valve is opened toallow a fluid flow, the circumstances changes and, although the total energy mustremain the same, it now consists of static pressure + dynamic pressure. As thevelocity V increases, so dynamic pressure increases and the static pressure isreduced.

1.7.4 TOTAL ENERGY.

Total energy can be measured as a ram pressure and is usually called the ‘totalhead’ or pitot pressure (PT). It is measured by placing a ram tube in the fluid flow.The ram tube must be parallel to the flow with its open end facing the flow. Agauge connected into such a tube always records the total head (pitot) pressureregardless of the rate of flow, refer to Fig 1.9.

Duct with Flow Control ValveFigure 1.8.



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In a situation where there is a no fluid flow, the static pressure (PS) gauge, and thetotal head pressure (PT) gauge will show the same value, but when there is a fluidflow, the total pressure reading remains the same although the static pressuredrops.

1.8 CONTINUITY EQUATION AND BERNOULLI’S THEOREM

1.8.1 INCOMPRESSIBLE FLUID.

The combined effect of the continuity equation and Bernoulli’s theorem producesthe effects shown, when a steady flow of incompressible fluid flows through a ductof varying cross sectional area (Fig 1.10.).

The effects of a steady flow of incompressible fluid flows through a duct of varyingcross sectional area shows:

Illustration of Pitot and Static PressuresFigure 1.9.

Duct of Varying Cross Sectional AreaFigure 1.10.



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Mass flow remains constant as cross-sectional area of duct (and velocity)change.

Total pressure remains constant, but static pressure (PS) changes as area(and velocity) change.

1.8.1.1 Compressibi lity Fluid (Atmosphere).

Compressible fluid flow refers to the air flow through a gas turbine engine and,because the air is compressible, flow at subsonic speeds causes a change in thedensity of the air as it progresses through the engine.

The air entering the duct at section A (Fig 1.11), consists of air at pressure (P1)

and velocity (V1); then as the air enters the increased area of the duct at B it willspread out to fill the increased area and this will cause the air flow to slow down(continuity equation) and give a change in velocity to V2. The static pressure ofthe air will increase (Bernoulli’s theorem) to become P2 in the wider section of theduct and, because air is compressible, the air density will increase as it iscompresses by the rise in pressure in section B of the duct.

Airflow Through a Duct Section

Figure 1.11.



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1.8.1.2 Diffuser action.The flare, which increases the area of the duct, is known as a diffuser (Fig1.12.)and its shape determines the rate of compression and the amount by whichthe air is compressed. For best results, the airflow must remain smooth and,because of this, a most important design feature is the angle of divergence. Whenair is compressed by this process it is called subsonic diffusion and it is a principlethat is used extensively in jet engine design.

1.8.2 GAS LAWS

In addition to the preceding information, the following gas laws are closely relatedto the function of a gas turbine engine:

Boyle’s Law. This law is related to temperature and pressure of a gas. Itstates that if the temperature T remains constant, the volume V of a given massvaries inversely as the pressure P exerted upon it (ie. PV = Constant).

Charles’ Law. This law states that the volume V of a given mass of gasincreases by 1/273 of its volume at 0°C for a rise of 1°C when the pressure P of

the gas is kept constant. These laws are now combined in what is called theideal gas law. It gives the relationship:

PV = RT where: P = pressure

V = volume

R = a constant

T = absolute temperature in K

Diffuser SectionFigure 1.12.



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1.9 SUBSONIC AIRFLOW THROUGH DIVERGENT AND CONVERGENTDUCTS

1.9.1 DIVERGENT DUCT

A divergent duct widens out as the airflow progresses through it. At subsonicspeeds the effect of this kind of duct is to decrease the velocity and increase thepressure and temperature of the air passing through it.

1.9.2 CONVERGENT DUCT

A convergent duct is such that the space inside reduces as the airflow progressesthrough it. At subsonic speeds the effect of this kind of duct is to increase thevelocity and decreases the pressure and temperature of the air passing through it.

Divergent Duct.Figure 1.13.

Convergent Duct.Figure 1.14.



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1.10 SONIC AIRFLOW THOUGH DIVERGENT AND CONVERGENT DUCTS

When a flow of fluid (i.e. gas) flows at sonic speed through a convergent duct ashock wave forms at the exit area of the duct - The exit area is said to be choked.The shock wave forms a restriction to the fluid and pressure will increase,temperature will increase and velocity will decrease.

When a gas flow reaches sonic velocity in a convergent duct the nozzle will chokeand the pressure will increase. To prevent a pressure rise that would eventually

prevent a 'fluid' flow and completely choke the duct a divergent section is addedmaking the duct convergent/divergent (Con/DI). The pressure of gas released intothe divergent section of the nozzle causes the velocity of the 'fluid' to increase,pressure to decrease, and therefore temperature to decrease. Gas pressure actson the walls of the divergent section, this pressure gives additional thrust that isknown as pressure thrust.

Airflow Through a Con-Di Nozzle or Venturi.Figure 1.15.

A Con-Di NozzleFigure 1.14.



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1.11 THE WORKING CYCLE ON A PRESSURE VOLUME DIAGRAM

Air is drawn from the atmosphere (Ambient Air) into the compressor. Thecompressor raises the pressure of the air (A to B) on diagram. If the pressure ofthe air is increased the volume is decreased. The air passes to the combustionsystem and heat is added by burning fuel with a proportion of the air. From thediagram (B to C) it is seen that combustion takes place at constant pressure so thegas turbine working cycle is known as the constant pressure cycle. In thecombustion system the air expands rearwards and the volume of the gasincreases and the gas kinetic energy increases. The gas flow passes to theturbine section to drive the turbine (s), energy is extracted and the pressure

decreases. The gas passes via an exhaust unit to the propelling nozzle whichforms a convergent duct. The velocity of the gas increases. The reaction to thehigh velocity jet produces thrust (C to D on diagram).

Changes in Temperature, Pressure and Velocity and the Brayton CycleFigure 1.16.



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1.12 ENGINE CONFIGURATIONS.

There are two main types of gas turbine engines: Reaction engines, which derive their thrust by jet reaction

Power engines, which provide a mechanical output to drive another device.

1.12.1 REACTION ENGINES

These can be divided into several categories.

a. Turbojet engines. The turbojet was the first type of jet engine developed. In thisengine all the air passes through the core engine (i.e. the compressor,

combustor and turbine). The engine may be single shaft as in the Avon engine,or twin shafted as in the Olympus 593 fitted to Concorde.

These engines are noisy and are not the most fuel efficient for normal use,however for high altitude high speed flight they are in a class of their own.

Turbo jet Engines.Figure 1.17.



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b. Low and Medium By-pass or turbofan engines. These engines will have two orthree shafts. The Low Pressure (LP) shaft drives a larger diameter compressor.Some of the air produced by-passes the core engine (hence the name) and isused to provide thrust. The core airflow provides power for the compressorsand thrust. These engine are quieter than turbojets and more fuel efficient. TheSpey and Tay engines fall into this category.

The by-pass ratio is determined by the ratio of the air in flowing through the by-pass to the air passing through the core of the engine. Low by-pass less than2:1, medium by-pass 2:1 to 4:1, high by pass greater than 5:1.

c. High by-pass turbofan engines. These engines have very large fans driven bya relatively small core engine. Often the fan is geared to run at a lower speedthan the LP turbine, which gives the turbine mechanical advantage and alsoallows it to run at higher speed where it is more efficient. The ALF 502, RB211and the Trent engines are all high by-pass

High by-pass enginesare very fuel efficient,powerful and quiet.

These engines have avery large diameterwhich does give dragproblems, and are notsuitable for highspeed flight as theblade tips will suffercompressibilityproblems as theyapproach the speedof sound.

Low By-pass Twin Spool Engine (Spey)Figure 1.17.

A Three Spool High By-pass Engine (RB211)Figure 1.18.



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1.12.2 POWER ENGINES

Power producing engines come in two main forms Turboprop and turboshaft.a. Turboprop Engines. Turboprop engines extract most of the energy from the

gas stream and convert it into rotational energy to drive a propeller. Theengines are either single or twin shaft and may be direct drive where the LP ormain shaft drive the propeller through a gearbox, or they may have a separatepower turbine to drive the propeller. Turboprop engines differ from high by-pass turbofans in that the propeller does not have an intake to slow andprepare the air before passing through it. The propeller therefore has to meetthe demands of airspeed etc. Examples of turboprops are the Dart, PW125and Tyne engines.

Turboprop EnginesFigure 1.19.



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b. Turboshaft Engines. These engines are used in helicopters. They sharemany of the attributes of turboprop engines, but are usually smaller. They do nothave propeller control systems built into the engine and usually do not have manyaccessories attached such as generators etc. as these are driven by the mainrotor gearbox. Modern turboshaft and turbo prop engines run at constant speedwhich tends to prolong the life of the engine and also means that they are moreefficient as the engine can run at its optimum speed all the time.

There are other types of engine such as ram jets, pulse jets, turbo-ram jet andturbo - rockets, but none of these are used commercially if at all.

Turboshaft Engine with Free power Turbine. (Gem)Figure 1.20.



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2 ENGINE PERFORMANCE

2.1 METHOD OF CALCULATING THE THRUST FORCES

The thrust forces or gas loads can be calculated for the engine, or for any flowsection of the engine, provided that the areas, pressures, velocities and mass floware known for both the inlet and outlet of the particular flow section.

The distribution of thrust forces shown in Fig 2.1. can be calculated by consideringeach component in turn and applying some simple calculations. The thrustproduced by the engine is mainly the product of the mass of air passing throughthe engine and the velocity increase imparted to it (ie. Newtons Second Law ofMotion), however the pressure difference between the inlet to and the outlet from

the particular flow section will have an effect on the overall thrust of the engine andmust be included in the calculation.

To calculate the resultant thrust for a particular flow section it is necessary tocalculate the total thrust at both inlet and outlet, the resultant thrust being thedifference between the two values obtained.

Thrust Distribution of a Typical Single Spool Axial Flow Engine.Figure 2.1.

TOTAL THRUST 11158 lbs

FORWARD GAS LOAD 57836 lbs REARWARD GAS LOAD 46678 lbs



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Calculation of the thrust is achieved using the following formula:

Thrust =g

WvP A J )(

Where A = Area of flow section in sq. in.

P = Pressure in lb. per sq. in.

W = Mass flow in lb. per sec.

VJ = Velocity of flow in feet per sec.

g = Gravitational constant 32.2 ft. per sec. per sec.

2.2 CALCULATING THE THRUST OF THE ENGINE

When applying the above method to calculate the individual thrust loads on thevarious components it is assumed that the engine is static. The effect of aircraftforward speed on the engine thrust will be dealt with later. In the followingcalculations ‘g’ is taken to be 32 for convenience.

Compressor casing

To obtain the thrust on the compressor casing, it is necessary to calculate theconditions at the inlet to the compressor and the conditions at the outlet from thecompressor. Since the pressure and the velocity at the inlet to the compressor arezero, it is only necessary to consider the force at the outlet from the compressor.

Therefore, given that the compressor –OUTLET Area (A) = 182 sq. in.

Pressure (P) = 94 lb. per sq. in. (gauge)

Velocity (v j) = 406 ft. per sec.

Mass flow (W) = 153 lb. per sec.

The thrust

= 0)( g

WvP A

j

= 032

406153)94182(

= 19,049lb. of thrust in a forward direction.



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Total Thrust of the Compressor.Figure 2.2.



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International Standard Atmosphere

Figure 2.3.



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Choked Nozzle

Considering the formula for thrust under “choked” nozzle conditions:

Thrust = 0( PP )A +g

Wv J

Where: P = PressureP = Ambient Pressure A = AreaW = Mass FlowV = Velocity

It can be seen that the thrust can be further affected by a change in the mass flowrate of air through the engine and by a change in jet velocity. An increase in massairflow may be obtained by using water injection to cool the air and increases in jetvelocity by using after-burning.

Changes in ambient pressure and temperature considerably influence the thrustof the engine. This is because of the way they affect the air density and hence themass of air entering the engine for a given engine rotational speed.

Thrust Correction - Turbojet

To enable the performance of similar engines to be compared when operatingunder different climatic conditions, or at different altitudes, correction factors mustbe applied to the calculations to return the observed values to those which would

be found under I.S.A. conditions. For example, the thrust correction for a turbo-jetengine is:

Thrust (lb) (corrected) = thrust (lb) (observed) xOP

30

Where P0 = atmospheric pressure in inches of mercury (in Hg)(observed)

30 = I.S.A. standard sea level pressure (in Hg)

Shaft Horsepower Correction - Turboprop

The observed performance of the turbo-propeller engine is also corrected to I.S.A.conditions, but due to the rating being in s.h.p. and not in pounds of thrust thefactors are different. For example, the correction for s.h.p. is:

S.h.p. (corrected) = s.h.p. (observed)OO T P

273

1527330

Where P0 = atmospheric pressure (in Hg) (observed)

T0 = atmospheric temperature in deg. C (observed)

30 = I.S.A. standard sea level pressure (in Hg)

273 + 15 = I.S.A. standard sea level temperature in deg. K273 + T0 = Atmospheric temperature in deg. K



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Equivalent Shaft Horsepower (EHP)

In practice there is always a certain amount of jet thrust in the total output of theturbo-propeller engine and this must be added to the s.h.p. The correction for jetthrust is the same as that specified earlier.

To distinguish between these two aspects of the power output, it is usual to referto them as s.h.p. and thrust horse-power (t.h.p.). The total equivalent horse-poweris denoted by t.e.h.p. (sometimes e.h.p.) and is the s.h.p. plus the s.h.p. equivalentto the net jet thrust. For estimation purposes it is taken that, under sea-level staticconditions, one s.h.p. is equivalent to approximately 2.6 lb. of jet thrust. Therefore:

t.e.h.p. = s.h.p.6.2

.lbthrust jet

The ratio of jet thrust to shaft power is influenced by many factors. For instance,the higher the aircraft operating speed the larger may be the required proportion oftotal output in the form of jet thrust. Alternatively, an extra turbine stage may berequired if more than a certain proportion of the total power is to be provided at theshaft. In general, turbo-propeller aircraft provide one pound of thrust for every 3.5h.p. to 5 h.p.

2.2.1 COMPARISON BETWEEN THRUST AND HORSE-POWER

Because the turbo-jet engine is rated in thrust and the turbo-propeller engine ins.h.p., no direct comparison between the two can be made without a power

conversion factor. However, since the turbo-propeller engine receives its thrustmainly from the propeller, a comparison can be made by converting the horse-power developed by the engine to thrust or the thrust developed by the turbo-jetengine to t.h.p.; that is, by converting work to force or force to work. For thispurpose, it is necessary to take into account the speed of the aircraft.

t.h.p. is expressed assec.550 per ft

FV

Where F = lb of thrust

V = aircraft speed (ft. per sec)

Since one horse-power is equal to 550ft.lb. per sec. and 550 ft. per sec. isequivalent to 375 miles per hour, it can be seen from the above formula that onelb. of thrust equals one t.h.p. at 375 m.p.h. It is also common to quote the speedin knots (nautical miles per hour); one knot is equal to 1.1515 m.p.h. or one poundof thrust is equal to one t.h.p. at 325 knots.

Thus if a turbo-jet engine produces 5,000 lb. of net thrust at an aircraft speed of

600 m.p.h. the t.h.p. would be 000,8375

600000,5



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However, if the same thrust was being produced by a turbo-propeller engine with apropeller efficiency of 55 percent at the same flight speed of 600 m.p.h., then the

t.h.p. would be: 545,1455

100000,8

Thus at 600 m.p.h. one lb. of thrust is the equivalent of about 3 t.h.p.

2.3 ENGINE THRUST IN FLIGHT

Since reference will be made to gross thrust, momentum drag and net thrust, it willbe helpful to define these terms:

Gross or total thrust is the product of the mass of air passing through the engineand the jet velocity at the propelling nozzle, expressed as:

0( PP )A +g

Wv J

The momentum drag is the drag due to the momentum of the air passing into the

engine relative to the aircraft velocity, expressed asg

WV where:

W = Mass flow in lb. per sec.

V = Velocity of aircraft in feet per sec.

G = Gravitational constant 32.2 ft. per sec. per sec.

The net thrust or resultant force acting on the aircraft in flight is the differencebetween the gross thrust and the momentum drag. From the definitions andformulae stated earlier under flight conditions, the net thrust of the engine,

simplifying, can be expressed as:

g

V V W APP

jo

All pressures are total pressures except P which is static pressure at the propellingnozzle

W = Mass of air passing through engine (lb. Per sec.)

VJ = Jet velocity at propelling nozzle (ft. per sec)

P = Static pressure across propelling nozzle (lb. Per sq. in)

PO = Atmospheric pressure (lb. Per sq. in)

A = Propelling nozzle area (sq. in)

V = Aircraft speed (ft. per sec.)

G = Gravitational constant 32.2

APPThrust essure

g

WV Thrust Momentum

g

wv APPThrust Gross

g

WV Drag Momentum

O

J

J o

)(Pr )(



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The Balance of Forces and Expression for Thrust and Momentum Drag.

Figure 2.4.

Graph of Thrust Against Forward Speed.Figure 2.5.



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2.3.1 EFFECT OF FORWARD SPEEDSince reference will be made to ‘ram ratio’ and Mach number, these terms aredefined as follows:

Ram ratio is the ratio of the total air pressure at the engine compressor entry tothe static air pressure at the air intake entry.

Mach number is an additional means of measuring speed and is defined as theratio of the speed of a body to the local speed of sound. Mach 1.0 thereforerepresents a speed equal to the local speed of sound.

From the thrust equation, it is apparent that if the jet velocity remains constant,

independent of aircraft speed, then as the aircraft speed increases the thrustwould decrease in direct proportion. However, due to the ‘ram ratio’ effect fromthe aircraft forward speed, extra air is taken into the engine so that the massairflow and also the jet velocity increase with aircraft speed. The effect of thistends to offset the extra intake momentum drag due to the forward speed so thatthe resultant loss of net thrust is partially recovered as the aircraft speedincreases. A typical curve illustrating this point is shown in the figure 2.5.Obviously, the ‘ram ratio’ effect, or the return obtained in terms of pressure rise atentry to the compressor in exchange for the unavoidable intake drag, is ofconsiderable importance to the turbo-jet engine, especially at high speeds. Abovespeeds of Mach 1.0, as a result of the formation of shock waves at the air intake,

this rate of pressure rise will rapidly decrease unless a suitably designed air intakeis provided; an efficient air intake is necessary to obtain maximum benefit from theram ratio effect.

As aircraft speeds increase into the supersonic region, the ram air temperaturerises rapidly consistent with the basic gas laws. This temperature rise affects thecompressor delivery air temperature proportionally and, in consequence, tomaintain the required thrust, the engine must be subjected to higher turbine entrytemperatures. Since the maximum permissible turbine entry temperature isdetermined by the temperature limitations of the turbine assembly, the choice ofturbine materials and the design of blades and stators to permit cooling are very

important.With an increase in forward speed, the increased mass airflow due to the ‘ramratio’ effect must be matched by the fuel flow and the result is an increase in fuelconsumption. Because the net thrust tends to decrease with forward speed, theend result is an increase in specific fuel consumption (s.f.c.), as shown by thecurves for a typical turbo-jet engine in the figure 2.6.

At high forward speeds at low altitudes, the ‘ram ratio’ effect causes very highstresses on the engine and, to prevent over-stressing, the fuel flow is automaticallyreduced to limit the engine speed and airflow.



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Effects of speed on Thrust and Fuel Consumption.Figure 2.6.



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2.3.2 EFFECT OF AFTERBURNING ON ENGINE THRUST At take-off conditions, the momentum drag of the airflow through the engine isnegligible, so that the gross thrust can be considered to be equal to the net thrust.If after-burning is selected, an increase in take-off thrust in the order of 30 percentis possible with the pure jet engine and considerably more with the by-passengine. This augmentation of basic thrust, is of greater advantage for certainspecific operating requirements.

Under flight conditions, however, this advantage is even greater, since themomentum drag is the same with or without after-burning and, due to the rameffect, better utilisation is made of every pound of air flowing through the engine.

2.3.3 EFFECT OF ALTITUDE

With increasing altitude the ambient air pressure and temperature are reduced.This affects the engine in two inter-related ways:-

The fall of pressure reduces the air density and hence the mass airflow into theengine for a given engine speed. This causes the thrust or s.h.p. to fall. The fuelcontrol system adjusts the fuel pump output to match the reduced mass airflow, somaintaining a constant engine speed.

The fall in air temperature increases the density of the air, so that the mass of airentering the compressor for a given engine speed is greater. This causes the

mass airflow to reduce at a lower rate and so compensates to some extent for theloss of thrust due to the fall in atmospheric pressure. At altitudes above 36,089feet and up to 65,617 feet, however, the temperature remains constant, and thethrust or s.h.p. is affected by pressure only.

Graphs showing the typical effect of altitude on thrust and fuel consumption areillustrated in Figure 2.7.



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2.3.4 EFFECT OF TEMPERATUREOn a cold day the density of the air increases so that the mass of air entering thecompressor for a given engine speed is greater, hence the thrust or s.h.p. ishigher. The denser air does, however, increase the power required to drive thecompressor or compressors; thus the engine will require more fuel to maintain thesame engine speed or will run at a reduced engine speed if no increase in fuel isavailable.

On a hot day the density of the air decreases, thus reducing the mass of airentering the compressor and, consequently, the thrust of the engine for a givenr.p.m. Because less power will be required to drive the compressor, the fuel

control system reduces the fuel flow to maintain a constant engine rotationalspeed or turbine entry temperature, as appropriate; however, because of the

decrease in air density, the thrust will be lower. At a temperature of 45C,depending on the type of engine, a thrust loss of up to 20 percent may beexperienced. This means that some sort of thrust augmentation, such as waterinjection, may be required.

The fuel control system, controls the fuel flow so that the maximum fuel supply isheld practically constant at low air temperature conditions, whereupon the enginespeed falls but, because of the increased mass airflow as a result of the increasein air density, the thrust remains the same. For example, the combined

acceleration and speed control (CASC) fuel system schedules fuel flow to maintaina constant engine r.p.m., hence thrust increases as air temperature decreasesuntil, at a predetermined compressor delivery pressure, the fuel flow isautomatically controlled to maintain a constant compressor delivery pressure and,therefore, thrust, Figure 2.8. illustrates this for a twin-spool engine where thecontrolled engine r.p.m. is high pressure compressor speed and the compressordelivery pressure is expressed as P3. It will also be apparent from this graph thatthe low pressure compressor speed is always less than its limiting maximum andthat the difference in the two speeds is reduced by a decrease in ambient airtemperature. To prevent the L.P. compressor overspeeding, fuel flow is alsocontrolled by an L.P. governor which, in this case, takes a passive role.

The Effect of AirTemperature ona Typical TwinSpool Engine

Figure 2.8.



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2.4 PROPULSIVE EFFICIENCY

Performance of the jet engine is not only concerned with the thrust produced, butalso with the efficient conversion of the heat energy of the fuel into kinetic energy,as represented by the jet velocity, and the best use of this velocity to propel theaircraft forward, ie. the efficiency of the propulsive system.

The efficiency of conversion of fuel energy to kinetic energy is termed thermal orinternal efficiency and, like all heat engines, is controlled by the cycle pressureratio and combustion temperature. Unfortunately this temperature is limited by thethermal and mechanical stresses that can be tolerated by the turbine. Thedevelopment of new materials and techniques to minimise these limitations iscontinually being pursued.

The efficiency of conversion of kinetic energy to propulsive work is termed thepropulsive or external efficiency and this is affected by the amount of kineticenergy wasted by the propelling mechanism. Waste energy dissipated in the jetwake, which represents a loss, can be expressed as

g

V vW j

2

)( 2where (vJ - V) is the waste velocity.

It is therefore apparent that at the aircraft lower speed range the pure jet streamwastes considerably more energy than a propeller system and consequently isless efficient over this range. However, this factor changes as aircraft speed

increases, because although the jet stream continues to issue at a high velocityfrom the engine, its velocity relative to the surrounding atmosphere is reducedand, in consequence, the waste energy loss is reduced.

Efficiency Plots of Differing Types of Engine to AirspeedFigure 2.9.



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2.5 FUEL CONSUMPTION AND POWER TO WEIGHT RELATIONSHIP

Primary engine design considerations, particularly for commercial transport duty,are those of low specific fuel consumption and weight. Considerable improvementhas been achieved by use of the by-pass principle and by advanced mechanicaland aerodynamic features and the use of improved materials. With the trendtowards higher by-pass ratios, in the range of 15:1, the triple-spool and contra-rotating rear fan engines allow the pressure and by-pass ratios to be achieved withshort rotors, using fewer compressor stages, resulting in a lighter and morecompact engine.

S.f.c. is directly related to the thermal and propulsive efficiencies; that is, theoverall efficiency of the engine. Theoretically, high thermal efficiency requires highpressures which in practice also means high turbine entry temperatures. In a pureturbo-jet engine this high temperature would result in a high jet velocity andconsequently lower the propulsive efficiency. However, by using the by-passprinciple, high thermal and propulsive efficiencies can be effectively combined byby-passing a proportion of the L.P. compressor or fan delivery air to lower themean jet temperature and velocity. With advanced technology engines of high by-pass and overall pressure ratios, a further pronounced improvement in s.f.c. isobtained.

The turbines of pure jet engines are heavy because they deal with the total airflow,whereas the turbines of by-pass engines deal only with part of the flow; thus theH.P. compressor, combustion chambers and turbines, can be scaled down. The

increased power per lb. of air at the turbines, to take advantage of their fullcapacity, is obtained by the increase in pressure ratio and turbine entrytemperature. It is clear that the by-pass engine is lighter, because not only has thediameter of the high pressure rotating assemblies been reduced, but the engine isshorter for a given power output. With a low by-pass ratio engine, the weightreduction compared with a pure jet engine is in the order of 20 per cent for thesame air mass flow.

With a high by-pass ratio engine of the triple-spool configuration, a furthersignificant improvement in specific weight is obtained. This is derived mainly fromadvanced mechanical and aerodynamic design, which in addition to permitting a

significant reduction in the total number of parts, enables rotating assemblies to bemore effectively matched and to work closer to optimum conditions, thusminimising the number of compressor and turbine stages for a given duty. Theuse of higher strength lightweight materials is also a contributory factor.

For a given mass flow, less thrust is produced by the by-pass engine due to thelower exit velocity. Thus, to obtain the same thrust, the by-pass engine must bescaled to pass a larger total mass airflow than the pure turbo-jet engine. Theweight of the engine, however, is still less because of the reduced size of the H.P.section of the engine. Therefore, in addition to the reduced specific fuelconsumption, an improvement in the power-to-weight ratio is obtained.



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2.6 SPECIFIC FUEL CONSUMPTION

When comparing engine performance, one of the most important considerations ishow efficiently the power is produced. The amount of fuel consumed to produce agiven horsepower lbs. thrust is known as “specific fuel consumption” or SFC. Atypical aircraft fuel system measures the volume of fuel consumed. This isdisplayed in pounds per hour or PPH. To calculate fuel flow, specific fuelconsumption found on the customer data sheet, is multiplied by the horsepowerlbs. thrust produced.

2.6.1 SPECIFIC FUEL CONSUMPTION – DEFINITION

SFC = SPECIFIC FUEL CONSUMPTION is defined as the lbs of fuel used perHP/lbs of thrust per hour

2.7 FLAT RATING

“Flat rating” is used by aircraft manufacturers when they select an engine that hasa capability greater than the requirements of the aircraft. They then limit the poweroutput of the engine. There are three distinct benefits derived from flat rating.One is the engine will have the ability to make take-off power at lower turbinetemperatures over a wide range of outside air temperatures and pressurealtitudes. Performance at altitude will be greatly enhanced. These two benefitsresult in the third benefit, longer engine life. A fourth benefit available on someengines is, a reserve of power which can be used to boost performance in anemergency ie. Loss of an engine during take - off.

2.8 PERFORMANCE RATINGS

In the chart, performance ratings are compared on –1 through –12 engines.Notice the modifiers on the –1, -5, -6, -8 and –10 engines. These temperaturesrepresent the effects of flat rating engines. Each engine will make take-off powerbelow their turbine temperature limits to the ambient temperatures indicated.Engines that are not flat rated, such as the –3 or –11, would be unable to maketake-off power below their turbine temperature limits when operating in conditions

above 59F outside air temperatures.



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3 INLET

3.1 INTRODUCTION

An air intake should deliver air to the engine compressor with a minimum loss ofenergy and at a uniform pressure under all engine operating conditions. The inletduct is built in the shape of a subsonic divergent diffuser, so that the kinetic energyof the rapidly moving air can be converted into a ram pressure rise within the duct.This condition is referred to as “Ram Recovery”.

3.2 RAM COMPRESSION

The degree of Ram Compression depends upon the following:-

i. Frictional losses at those surfaces ahead of the intake entry which are

“wetted” by the intake airflow.ii. Frictional losses at the intake duct walls.

iii. Turbulence losses due to accessories or structural members located in theintake.

iv. Aircraft speed.

v. In a turbo-prop, drag and turbulence losses due to the prop blades andspinner.

Ram compression causes a re-distribution in the forms of energy existing in theair-stream. As the air in the intake is slowed up in endeavouring to pass into andthrough the compressor element against the air of increasing pressure and densitywhich exists therein so the kinetic energy of the air in the intake decreases. Thisis accompanied by a corresponding increase in its pressure and internal energiesand consequently compression of the air-stream is achieved within the intake, thusconverting the unfavourable intake lip conditions into the compressor inletrequirements.

Although ram compression improves the performance of the engine, it must berealised that during the process there is a drag force on the engine and hence theaircraft. This drag must be accepted since it is a penalty inherent in a ramcompression process. (The added thrust more than makes up for this drag).

3.2.1 IMPORTANCE OF RAM COMPRESSION

At subsonic flight speeds, the ram pressure ratio is apparently quite small, say1.33: 1 at 0.8M. Nevertheless, since the pressure rise due to ram compression ismultiplied by the pressure ratio of the compressor, the ram pressure rise becomessignificant even at subsonic speeds.

Furthermore, the greater the forward speed of the aircraft becomes, the moresignificant is the ram compression; e.g. at 1.5M the ram pressure ratio may beabout 3.5 : 1, and at 2.5M about 8 : 1.



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3.3 TYPES OF AIR INTAKES

3.3.1 PITOT INTAKESThis intake is suitable for subsonic or low supersonic speeds. Examples, 707,747, A300B, Tristar, etc. The intake is usually short and is very efficient becausethe duct inlet is located directly ahead of the engine compressor. As the ductlength increases, the risk of small airflow disturbances and pressure drop isincreased. This inlet makes maximum use of ram effect until sonic speed isapproached when efficiency falls due to shock wave formation at the intake lip.Pitot inlets can however suffer from inlet turbulences at high angles of attackand/or at low speeds.

Pitot Type Intakes.

Figure 3.1.



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The pitot type intake can be used for engines that are mounted in pods or in thewings although the latter sometimes requires a departure from the circular crosssection due to the wing thickness.

3.3.2 DIVIDED ENTRANCE DUCT

On a single engine aircraft with fuselage mounted engines, either a wing root inletor a side scoop inlet may be used. The wing root inlet presents a problem todesigners in the forming of the curvature necessary to deliver the air to the enginecompressor. The side scoop inlet is placed as far forward of the compressor aspossible to approach the straight line effect of the single inlet. Both types sufferfaults, in a yaw or turn, a loss of ram pressure occurs on one side of the intake andseparated, turbulent boundary layer air is fed to the engine compressor.

Wing Leading Edge IntakesFigure 3.2

Divided Intakes.Figure 3.3.



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3.3.3 SUPERSONIC INTAKES

At supersonic speeds, the pitot type of air intake is unsuitable due to the severityof shock waves which form and progressively reduce the intake efficiency asspeed increases. To overcome this problem the compression intake wasdesigned.

This type of intake produces a series of mild shock waves without reducing the

intake efficiency, as the aircraft speed increases, so also does the intakecompression ratio. At high mach numbers it becomes necessary to have an airintake which has a variable thrust area and spill doors to control the column of air.

3.4 IDEAL INTAKE CONDITIONS

For air to flow smoothly through a compressor, its velocity should be about 0.5mach at the compressor inlet; this includes aircraft flying faster than the speed ofsound. Hence intakes are designed to decelerate the free stream airflow to thiscondition over the range of aircraft speeds. Intakes should also convert the kineticenergy into pressure energy without undue shock or energy loss. This meansthat the ideal compressor inlet pressure should be the same as the total head

pressure at the intake lip.

(Total head pressure = stagnation pressure, ie. static and dynamic pressure).

Supersonic Intakes.Figure 3.4.



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Intake EfficiencyThe magnitude of the losses occurring in an intake during ram compression aremeasured by means of the intake efficiency. Typical optimum efficiencies of somecommon types of intake, at subsonic speeds assuming straight-through flow, are:

a Turbo-jet engine Pitot 99 to 96%

Wing root 95 to 87%

Side 89 to 80%

b Turbo-prop engine Annular 82 to 74% (DART)

In cases where the direction of flow of the air is reversed within the intake, thesevalues are reduced by about 10%.

3.5 INTAKE ANTI-ICING

Operations of present day aircraft necessitates flying in all weather conditions plusthe fact that high velocity air induced into the intakes means a provision must bemade for ice protection. There are three systems of thermal anti-icing; hot air, hotoil or electrical There is, however, one disadvantage and that is the loss ofengine power. This loss must be corrected for on ground runs and power checks.

3.5.1 ENGINE HOT AIR ANTI-ICING

The hot air system provides surface heating of the engine and/or power plantwhere ice is likely to form. The affected parts are the engine intake, the intakeguide vanes, the nose cone, the leading edge of the nose cowl and, sometimes,the front stage of the compressor stator blades. The protection of rotor blades israrely necessary, because any ice accretions are dispersed by centrifugal action.

The hot air for the anti-icing system is usually taken from the latter stages of theHP compressor and externally ducted, through pressure regulation valves, to theparts requiring protection. When the nose cowl requires protection, hot airexhausting from the air intake manifold may be collected and ducted to the nosecowl. Exhaust outlets are provided to allow the air to pass into the compressorintake or vent to atmosphere, thus maintaining a flow of air through the system.



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Hot Air Anti-Icing.Figure 3.5.



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3.5.2 ENGINE ELECTRICAL ANTI-ICINGThere are two methods of electrical anti-icing:

1. Spray mat

2. Heater mats.

3.5.2.1 Spray Mat

The spray mat is so called because the conductor element is sprayed onto thebase insulator to protect the spray mat from damage. An outer coating is sprayedon, sometimes called “Stone Guard” or “Erocoat”.

Spraymat Construction.Figure 3.6.



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3.5.2.2 Heater Mats

Heater mats differ in design and construction according to their purpose andenvironment. The latest mats have elements which are made from a range ofalloys woven in continuous filament glass yarn. Other elements are made from

nickel chrome foil. The insulating material is usually polytetrafluoroethylene(PTFE) and the electrical control may be continuous or intermittent.

3.5.3 OIL ANTI-ICE

Oil anti-ice supplements the other two systems (hot air/electrical) and will alsoassist in cooling the oil system.

Heater Mat Construction.Figure 3.7.



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Hot Oil Anti-IceFigure 3.8.



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Intentionally Blank



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A Double Entry Centrifugal CompressorFigure 4.2.



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4.2.1 OPERATION

The centrifugal impeller is rotated at high speed by the turbine and centrifugalaction causes the air between the impeller vanes to accelerate radially outwardsuntil it is thrown off at the tip into the diffuser. The radial movement of the airacross the impeller, from eye to tip, causes a drop in air pressure at the eye andthe faster the impeller is turning, the lower the pressure at the eye becomes. Thelow pressure existing at the eye of the revolving impeller induces a continuous flowof air through the engine intake and into the eye of the impeller. The air, in turn, isaccelerated across the impeller and passed into the diffuser. The kinetic energy inthe air is then converted to pressure energy ready to enter the combustionchamber. The action of the diffuser is illustrated in figure 4.3.



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Centrifugal Compressor Function.Figure 4.3.

VANELESSSPACE



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The final volume and mass airflow delivered by the centrifugal compressor isdependent on:

a Pressure ratio

b Operating RPM

c Diameter of the impeller

NOTE: This is assuming a constant air density at the inlet of the compressor.

4.2.1.1 Pressure Ratio

The ratio of the inlet pressure to outlet pressure of the compressor is called

pressure ratio. The higher the pressure of the air the more efficiently the thrust willbe produced with a corresponding improvement to the fuel economy of the engine.

The maximum pressure ratio normally obtainable from a single stage centrifugalcompressor is approximately 5:1 and from a two stage, approximately 8:1.Designof the more modern centrifugal compressors sees them approaching pressureratios of 15:1.

4.2.1.2 Diameter of Impeller

A large impeller will deliver a greater mass of air than a small impeller, however alarge diameter compressor leads to an increase in the frontal area of the enginecausing excess drag forces on the aircraft.

4.3 THE AXIAL FLOW COMPRESSOR

The axial flow compressor is by far the most popular type of compressor and,although it is more difficult to manufacture, it is a more efficient compressor.Handling a larger mass of air for any given diameter, it produces more power; andbecause the compression ratio is high – at least 9:1 and, it can be very muchhigher – it is a more economical engine. The airflow through the engine is parallelwith the axis, hence the name ‘axial flow compressor’.

The compressor consists of a single or multi-rotor assembly that carries blades ofaerofoil section; it is mounted in a casing, which also houses the stator blades.

The axial flow compressor increases the pressure of the air gradually (byapproximately 1.2:1 per stage) over a number of ‘stages’, each stage comprisingof a row of ‘rotor blades’, followed by a row of ‘stator blades’. Both the rotor andstator blades are of aerofoil section and form divergent passageways betweenadjacent blades of the same row. Figure 4.4 refers.



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Axial Compressor Layouts.Figure 4.6.



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The engine rotor assembly forms a hollow “drum” and is supported in ball androller bearings and coupled to a turbine shaft. The rotor discs make up the drumand the rotor blades are attached as shown in the figure. On some smallerengines it becomes difficult to design a practical fixing, this is overcome bydesigning and producing blades integral with the disc and is called a “BLISK”.

Compressor Blade AttachmentFigure 4.9



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1 Extension Shaft Drive Stub 2 1st

Stage Disk3 Balance Weight 4 1st Stage Rotor Blades

5 Shroud Rings 6 7th Stage Rotor Blades

7 Air Inlet to Rotor Drum 8 1st Stage Blade Locking Strips

9 Front Main Bearing Housing

Axial Compressor Rotor Details.Figure 4.10.



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Axial Compressor Stator DetailsFigure 4.11



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The mass and final volume of the airflow delivered by the compressor isdependent on:

a. Pressure Ratio. Dependent on the number of stages employed. Axial flowcompressors can achieve a much higher value than centrifugal.

b. Diameter . For a similar mass flow capability, the axial flow compressor can bemade smaller in diameter than the centrifugal type.

c. Operating RPM. As with the centrifugal type, the RPM and hence the massflow, is controlled by varying the amount of fuel delivered to the combustionsystem, but because of the way that the pressure rise takes place, themaximum pressure ratio in an axial flow compressor is achieved at a lowerRPM, than in a centrifugal compressor.

4.4 COMPRESSOR STALL AND SURGE

‘Surge’ can occur in both centrifugal and axial flow compressors and is thereversal of the airflow in the compressor. It is a very undesirable condition, whichcan rapidly cause serious damage to the engine.

In an axial flow compressor, ‘surge’ is nearly always preceded by stalling of someof the compressor blades. An aerofoil is said to be in a stalled condition when theairflow over its surface has broken down and no lift is being produced. If a row ofcompressor blades stall, then they will not be able to pass the airflow rearwards to

the next stage and the airflow to the combustion chamber will ultimately stop.The lack of rearward airflow will allow the air in the combustion chamber to flowforward into the compressor until it reaches the row of stalled blades. Then aviolent backwards and forwards oscillation of the airflow is likely to occur, whichcan rapidly cause extensive damage to the compressor blades and also over-heating of the combustion and turbine assemblies.

Stalling of the compressor blades can occur for various reasons and to appreciatehow the condition comes about, a review of aerofoil theory and its application tothe compressor is required.

4.4.1 AIRFLOW CONTROL SYSTEM PRINCIPLES

4.4.1.1 Compressor Stall and Surge

For any given engine there is only one set of conditions, mass flow, pressure ratioand rpm, at which all the compressor components are operating at their optimumeffect. Compressors are designed to be most efficient in the higher rpm range ofoperation. The point at which the compressor reaches its maximum efficiency isknown as the DESIGN POINT. Under design conditions the compressor produces

a given compression ratio (ie.1

2

Volume

Volume) and the axial velocity (average velocity)

of the gas remains approximately constant from the front to the rear of the

compressor.



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The Angle of Attack of the airflow to the compressor aerofoil blades will be at itsoptimum. This is the design condition and the compressor is operating at itsoptimum performance. Although compression ratio varies with rpm it is notproportional to rpm. This fact emerges due to the fixed blade angles, which canonly be correct at the design point. To illustrate this fact, refer to the diagramshowing rpm and compression ratio. Consider a compressor running at 8,000 rpmand its compression ratio is 10:1. Let us say that the volume of air entering thecompressor is 100cm3. The volume of the air passing through the fixed outletannulus of the compressor will be 10cm3.

Graph of Compression Ratio to RPM.

Figure 4.12.

Compressor R.P.M = 8,000 Compressor R.P.M. = 4,000Compression Ratio = 10:1 Compression Ration = 4:1

Volume of gas (V1) = 100cm3 Volume of gas (V1) = 50cm3

Volume of gas (V2) = 10cm3 Volume of gas (V2) = 12.5cm3

C O M P R E S S I O N

R A T I O

10:1

4:1

4000 8000

RPM



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Now consider the same compressor operating at 4,000 rpm, the volume of airentering the compressor will be halved, eg. 50cm3 there will also be a reduction incompression ratio to 4:1. Therefore the volume of air passing through thecompressor fixed outlet annulus will be 12.5cm3. The following conditions willoccur:

a. Axial velocity will increase as it moves towards the rear stages relative to thefront Low pressure stages.

b…Since all stages are rotating at the same speed, there will be a NEGATIVEangle of attack at the rear high pressure stages and a POSITIVE angle of attack atthe front low pressure stages.

Effect of Velocity on Blade Angle.

Figure 4.13.

Front Rear



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Due to the increased velocity at the rear of the compressor, the outlet of thecompressor will choke as the airflow reaches sonic velocity. At this point there willbe a dramatic reduction in axial velocity resulting in the front compressor bladesstalling. The end result will be compressor surge. To overcome the problem, ableed valve is normally fitted in an intermediate stage of the compressor to bleedoff the excess volume of air. This relieves the rear stages of the excess aircausing choking while inducing an increased axial airflow through the early stagesof the compressor, thus establishing conditions which are not conducive of stalland surge. Unfortunately this bleed valve does not completely cure the problem ofstall as far as the first rotor stages are concerned and stall is still likely to occur.The blades stall when the angle of attack increases to too large a value. To

overcome this problem, inlet guide vanes are used to pre-swirl the air onto therotor blades. The effect of pre-swirling the air alters the angle of attack from alarge value to the correct angle of attack. See figure 4.14.

Effect of Variable Guide Vane on Compressor StageFigure 4.14



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4.4.3 EFFECT OF TEMPERATURE ON THE OPERATING POINT OF THE AIRFLOWCONTROL SYSTEM

A change in temperature will affect mass airflow, compressor pressure ratio fuelflow and engine performance. The effect of a reduced temperature on thecompressor at a fixed rpm being that the performance is comparable with that at ahigher rpm at STANDARD TEMPERATURE.

Consider an engine running at 10,000 rpm, the temperature of the day is 2ºC. Ifthis is corrected for standard conditions (ISA 15ºC) the corrected rpm will be10,235 see below.

Observed rpm = 10,000 rpm

Corrected rpm =

N

Where =K inISA

K inambientT=

15273

2273

corrected rpm =

288

275

000,10

=977.0

000,10

Corrected rpm = 10,235

From the above it is clear that temperature has an effect on the compressors massflow rate. This is compounded further by the effect that temperature has a directeffect on the speed of sound and hence when the compressor chokes.

It must be understood that if the engine is running at a fixed rpm and thetemperature of the air is altered, the actual rpm of the compressor will beunaffected. However, the temperature change will affect the mach number of

mass airflow and it is the speed of the compressor relative to the speed of theairflow (ie. Mach. Number) which is the critical factor. A decrease in temperaturewill raise the mach. Number. The mach. Number is the:

SOUNDOF SPEED LOCAL

OBJECT THE OF SPEED

ISAinK



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The speed of the object is the compressor blade, if as previously stated, the mach.Number is raised with a decrease in temperature, the ‘fixed’ blade speed relativeto the speed of the air, will be increased. To cater for this situation the operatingpoint at which the variable inlet guide vanes move will have to be altered forvarying air temperatures. To achieve this the actuator or ram of an airflow controlsystem is temperature compensated. On a ‘cold’ day, the variable inlet guidevanes will operate earlier than on a ‘warm’ day.

At a temperature of +60F Local speed of sound is Mach 0.9 , no need for theVIGV’s as the compressor out let is not choked.

At a temperature of –40°F Local speed of sound is Mach 1.0, the compressoroutlet is choked, the first stages may stall, VIGV’smust start to open.

Variation of Mach Number with Temperature.Figure 4.16.



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4.5 AIR FLOW CONTROL SYSTEM – OPERATIONThe stages of the compressor are matched to give the highest efficiency in thespeed range maximum rev/min. To extend the range of smooth operation overlower engine speeds, variable-incidence intake guide vanes and/or an air bleedvalve are fitted. In the lower speed range the bleed valve opens to allow some ofthe air to escape from the rear stages of the compressor, thus restricting the massair flow through the later stages and preventing an unstable flow pattern.

When the bleed valve is open, the guide vanes if fitted are partially closed; athigher engine speeds, when the bleed valve is closed, the guide vanes if fittedmove progressively towards the open position. The vanes are operated by a

hydraulic ram which incorporates its own control mechanism and which receives asignal of engine speed in terms of hydraulic pressure from the engine speedgovernor in the fuel pump.

Combined Bleed Valve and Variable Guide Vane Operating System.

Figure 4.17.



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Intake Guide Vane Ram Setting Curve.Figure 4.18.



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Intake Guide Vane Ram Setting Curve.

Figure 4.18.

Air Bleed ValveFigure 4.19.

Variable Guide Vane Hydraulic

Figure 4.20.



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To further improve airflow control, some engines will adopt a system of VariableStator Vanes (VSV’s) as well as Variable Inlet Guide Vanes (VIGV’s) figure 4.21.

Variable IGV and Stator Vanes.Figure 4.21.



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Inlet Guide Vane and Variable stator Blade Linkwork.Figure 4.22.



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4.6 AEROFOIL THEORY AND THE AXIAL FLOW COMPRESSOR(CONTINUED)

The blades of the axial flow compressor are aerofoils and as such behave in asimilar way to aircraft mainplanes and propeller blades. The airflow across theirsurfaces produces lift and the amount of lift produced by an aerofoil depends on:

a Its shape, area and smoothness of its surface.

b the speed of airflow over the aerofoil.

c the angle at which the aerofoil meets the air.

Once manufactured, their area and shape will remain the same unless they aredamaged in any way. Assuming the blades are in good condition, the variables

will be the speed of the airflow and the angle at which the blades meet the air(angle of attack).

4.6.1 SPEED OF AIRFLOW OVER BLADES

This will vary with the rpm of the compressor rotor. The faster the rotor turns, thenthe faster the air flows over the blades. This will result in an increase in the axialvelocity of the airflow through the compressor.

4.6.2 ANGLE OF ATTACK

This will vary with the combination of the rotational velocity of the blades and theaxial velocity of the airflow. In the normal course of events, the angle of attack

(VA) becomes progressively smaller as the compressor moves from a low rpm to ahigh rpm.(VT)

VT VTVT

VT

V A V A

V A

V A

Low R.P.M R.P.M Increasing High R.P.M

High angleof attack

Angle of attackdecreasing

Low angleof attack

Change of Angle of Attack Due to Increase in RPM.Fi ure 4.23.



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4.6.3 SOME IMPORTANT POINTS ABOUT ANGLE OF ATTACK

An aerofoil can only produce lift between certain limits of angle of attack. 0 -

approx. 15.

At very large angles of attack the airflow breaks down and the aerofoil stalls.

The greater the angle of attack (up to the stalling angle), the greater the lift and,also, the greater the drag. This means that a greater effort will be required tomove the aerofoil through the air.

All aerofoils have an ‘optimum’ angle of attack at which they produce most lift forthe least drag. (‘Lift/drag ratio’) [2-4°].

At High Angles of Attack the Blade Will Stall.Figure 4.25

Lift/drag Vectors for Different Angles of Attack.Figure 4.26.

Airflow Over an AerofoilFigure 4.24.



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4.7 APPLICATION TO THE AXIAL FLOW COMPRESSOR

In order for the compressor to deliver a high mass airflow for a minimum effortrequired to drive it, it is important that all the compressor blades are operatingclose to their optimum angle of attack at the designed optimum rpm of the engine.

This is achieved by setting the blades onto the rotor assembly at a large enoughangle so as to make allowance for the automatic reduction in angle of attack thatwill occur with increase in rpm.

4.7.1 COMPRESSOR RPM

An axial flow compressor is designed to operate at maximum speeds in the regionof 8000-10,000 rpm, depending on size. At this rpm the engine will be producing alarge amount of thrust and in order to vary the thrust it is necessary to vary the

compressor rpm.

When the compressor is operating at speeds below its designed rpm range, theaxial velocity of the airflow through the compressor will decrease which will causean increase in the angle of attack of the compressor blades. At low rpm, such asidling, the reduced axial velocity of the airflow may cause the angle of attack ofsome of the blades to increase beyond their stalling angle.

A slight amount of LP blade stalling during ‘off design’ conditions is to be expectedand only becomes a problem if a complete row of blades stall.

4.7.2 COMMON CAUSES OF COMPRESSOR STALL

Compressor stall normally occurs at low rpm and can be induced by:

a disturbance of smooth airflow due to damaged or dirty blades.

b disturbance of smooth airflow caused by damaged aircraft air intake.

c high combustion chamber pressure caused by over-fuelling during engineacceleration.

4.7.3 STAGGER ANGLE AND END BEND

The rotor blades are of airfoil section and usually designed to give a pressuregradient along their length to ensure that the air maintains a reasonably uniformaxial velocity. The higher pressure towards the tip balances out the centrifugalaction of the rotor on the airstream. To obtain these conditions, it is necessary to'twist' the blade from root to tip to give the correct angle of incidence at each point.

Air flowing through a compressor creates two boundary layers of slow to stagnantair on the inner and outer walls. In order to compensate for the slow air in theboundary layer a localised increase in blade camber both at the blade tip and roothas been introduced. The blade extremities appear as if formed by bending overeach corner, hence the term 'end-bend' Figure 4.27.

4.7.4 RECENT INNOVATIONS

The latest engines incorporate blades that have been designed and profiled using3-D design techniques. This produces blades, which are curved in 3 dimensions,which are more aerodynamically efficient. Figure 4.28.



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3-D BladesFigure 4. 28.

Stagger Angle and End BendFigure 4.27.



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4.8 AIRFLOW CONTROLThe higher the pressure ratio required from a compressor, the greater the numberof compressor stages needed. The more stages there are, the more difficultbecomes the problem of matching all the blades in both size and angle ofattachment to make the compressor operate satisfactorily over a wide range ofrpm.

In order to maintain the airflow stability and reduce the tendency of high pressureratio compressors to stall under certain conditions of aircraft flight and enginehandling, methods of airflow control have already been discussed.

4.9 AIR BLEED VALVES (SUMMARY)

The air bleed valve is operated automatically in response to signals of compressorrpm. It is in the open position below a certain critical rpm and bleeds air awayfrom the centre stages of the compressor, ducting it overboard to atmosphere.This has the effect of increasing the axial velocity of the airflow through the earlystages of the compressor, thereby reducing the angle of attack of the blades inthat area. This prevents the early stages of the compressor from passing more airto the rear stages than can be accommodated in the space available.

Above the critical rpm range the bleed valve is closed and all the air available fromthe compressor passes to the combustion system.

4.10 VARIABLE INTAKE GUIDE VANES (SUMMARY) All intake guide vanes give a certain amount of swirl to the incoming airflow. Theswirl is in the direction of rotation of the compressor and the amount of swirldetermines the angle of attack of the first stage rotor blades. The greater thedegree of swirl imported by the IGV’s then the smaller the resultant angle of attackof the first stage rotor blades.

Variable IGV’s present the air onto the first stage rotor blades with a maximumswirl angle during operation in the critical low rpm range and progressively reducethe degree of swirl in response to signals of compressor rpm. When operating athigh rpm the airflow enters the compressor more or less axially.

4.11 MULTI-SPOOL COMPRESSORS (SUMMARY)

Pressure ratios in excess of approximately 9:1 are best achieved by splitting thecompressor into two independent sections as shown in the figure 4.29.



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The total number of stages of compression is divided between two spools, eachspool being driven at a different speed by separate turbines. This eases theproblems of compressor blade matching and results in a very powerful, efficientand flexible engine.

4.12 COMPARING THE FEATURES OF CENTRIFUGAL AND AXIAL FLOWCOMPRESSORS

4.12.1 CENTRIFUGAL

Merits. Simplicity, cheaper, lighter, less prone to damage by FOD.

Not critical to surge and stall.

Will tolerate icing conditions.

Associated Problems

Max pressure ratios 4:1 or 5:1. (on early types)

Capacity limited by tip speed.

Larger diameter of engine which leads to more drag.

Severe directional changes of gas flow which leads to friction.

High specific fuel consumption.

4.12.2 AXIAL FLOW

Merits

High Pressure Ratio.

Low specific fuel consumption.

More capacity for development.

Greater axial thrust.

Associated Problems

Complex and expensive to produce.

Critical to stall/surge.

Twin Spool EngineFi ure 4.29.



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5 COMBUSTION SECTION

5.1 INTRODUCTION

The combustion chamber has the difficult task of burning large quantities of fuel,supplied through the fuel burners, with extensive volumes of air, supplied by thecompressor, and releasing the heat in such a manner that the air is expanded andaccelerated to give a smooth stream of uniformly heated gas at all conditionsrequired by the turbine. This task must be accomplished with the minimum loss inpressure and with the maximum heat release for the limited space available.

The amount of fuel added to the air will depend upon the maximum temperaturerise required and, as this is limited by the materials from which the turbine bladesand nozzles are made, the rise must be in the range of 700 to 1,200 deg.C.

Because the air is already heated by the work done during compression, thetemperature rise required at the combustion chamber may be between 500 and800 deg.C. Since the gas temperature required at the turbine varies with enginespeed, and in the case of the turbo-prop engine upon the power required, thecombustion chamber must also be capable of maintaining stable and efficientcombustion over a wide range of engine operating conditions.

Efficient combustion has become more and more important because of the rapidincrease in commercial aircraft traffic and the consequent increase in atmosphericpollution, which is seen by the general public as exhaust smoke.

5.2 COMBUSTION PROCESS

Air from the engine compressor enters the combustion chamber at a velocity up to500 feet per second, but because at this velocity the air speed is far too high forcombustion, the first thing that the chamber must do is to diffuse it, i.e. decelerateit and raise its static pressure. Because the speed of burning kerosene at normalmixture ratios is only a few feet per second, any fuel lit even in the diffused airstream, which now has a velocity of about 80 feet per second, would be blownaway. A region of low axial velocity has therefore to be created in the chamber, sothat the flame will remain alight throughout the range of engine operatingconditions.

In normal operation, the overall air/fuel ratio of a combustion chamber can varybetween 45:1 and 130:1. Kerosene, however, will only burn efficiently at, or closeto, a ratio of 15:1, so the fuel must be burned with only part of the air entering thechamber, in what is called a primary combustion zone. This is achieved by meansof a flame tube (combustion liner) that has various devices for metering the airflowdistribution along the chamber.



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Approximately 20 per cent of the air mass flow is taken in by the snout or entrysection. Immediately downstream of the snout are swirl vanes and a perforated

flare, through which air passes into the primary combustion zone. The swirling airinduces a flow upstream of the centre of the flame tube and promotes the desiredrecirculation. The air not picked up by the snout flows into the annular spacebetween the flame tube and the air casing.

Through the wall of the flame tube body, adjacent to the combustion zone, are aselected number of holes through which a further 20 per cent of the main flow ofair passes into the primary zone. The air from the swirl vanes and that from theprimary air holes interacts and creates a region of low velocity recirculation. Thistakes the form of a toroidal vortex similar to a smoke ring, and has the effect ofstabilising and anchoring the flame. The recirculating gases hasten the burning offreshly injected fuel droplets by rapidly bringing them to ignition temperature.

Typical Combustion ChamberFigure 5.1.



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It is arranged that the conical fuel spray from the burner intersects the recirculationvortex at its centre. This action, together with the general turbulence in theprimary zone, greatly assists in breaking up the fuel and mixing it with theincoming air.

The temperature of the combustion gases released by the combustion zone isabout 1,800 to 2,000 deg.C., which is far too hot for entry to the nozzle guidevanes of the turbine. The air not used for combustion, which amounts to about 60per cent of the total airflow, is therefore introduced progressively into the flametube. Approximately half of this is used to lower the gas temperature before itenters the turbine and the other half is used for cooling the walls of the flame tube.Combustion should be completed before the dilution air enters the flame tube,otherwise the incoming air will cool the flame and incomplete combustion will

result.

An electric spark from an igniter plug initiates combustion and the flame is thenself-sustaining.

The design of a combustion chamber and the method of adding the fuel may varyconsiderably, but the airflow distribution used to effect and maintain combustion isalways very similar to that described.

5.3 FUEL SUPPLY

So far little has been said of the way in which the fuel is supplied to the air stream.In general, however, two distinct principles are in use, one based on the injectionof a finely atomised spray into a recirculating air stream, and the other based onthe pre-vaporisation of the fuel before it enters the combustion zone.

Although the injection of fuel by atomiser jets is the most common method, someengines use the fuel vaporising principle. In this instance, the flame tube is of thesame general shape as for atomisation, but has no swirl vanes or flare. Theprimary airflow passes through holes in a baffle plate that supports a fuel feedtube.

Apportioning the Airflow

Figure 5.2



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The fuel is sprayed from the feed tube into vaporising tubes that are positionedinside the flame tube. These tubes bend through 180 degrees and, as they areheated by combustion, the fuel vaporises before passing forwards into the flametube. The primary airflow passes down the vaporising tubes with the fuel and alsothrough large (secondary) nozzles, which provide 'fans' of air to sweep the flamerearwards. Cooling and dilution air is metered into the flame tube in a manner

similar to the atomiser flame tube. Vaporisers require starter spray nozzles to setthe light up process in motion.

5.4 TYPES OF COMBUSTION CHAMBER

There are three main types of combustion chamber at present in use for gasturbine engines. These are the multiple chamber, the tubo-annular chamber andthe annular chamber.

5.4.1 MULTIPLE COMBUSTION CHAMBER

This type of combustion chamber is used on centrifugal compressor engines andthe earlier types of axial flow compressor engines. It is a direct development of

the early type of Whittle combustion chamber. The major difference is that theWhittle chamber had a reverse flow as this created a considerable pressure loss,the straight through multiple chamber was developed by Joseph Lucas Limited.

The chambers are disposed around the engine and compressor delivery air isdirected by ducts to pass into the individual chambers. Each chamber has aninner flame tube around which there is an air casing. The air passes through theflame tube snout and also between the tube and the outer casing as alreadydescribed.

The separate flame tubes are all interconnected. This allows each tube to operateat the same pressure and also allows combustion to propagate around the flame

tubes during engine starting.

A Vaporising Combustion Chamber.Figure 5.3.



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Multiple Combustion Chambers.Figure 5.4.



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5.4.2 TUBO-ANNULAR COMBUSTION CHAMBER(ALSO KNOWN AS CAN-ANNULAR OR CANNULAR.)

The tubo-annular combustion chamber is a combination of the multiple andannular types. A number of flame tubes are fitted inside a common air casing.The airflow is similar to that already described and this arrangement embodies theease of overhaul and testing of the multiple system with the compactness of theannular system.

Turbo-Annular Combustion SystemFigure 5.5.



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5.4.3 ANNULAR COMBUSTION CHAMBERThis type of combustion chamber consists of a single flame tube, completelyannular in form, which is contained in an inner and outer casing. The airflowthrough the flame tube is similar to that previously described, the chamber beingopen at the front to the compressor and at the rear to the turbine nozzles.

The main advantage of the annular chamber is that, for the same power output,the length of the chamber is only 75 per cent of that of a tubo-annular system foran engine of the same diameter, resulting in considerable saving of weight andproduction cost. Another advantage is that because interconnectors are notrequired, the propagation of combustion is improved.

In comparison with a tubo-annular combustion system, the wall area of acomparable annular chamber is much less; consequently, the amount of coolingair required to prevent the burning of the flame tube wall is less, by approximately15 per cent. This reduction in cooling air raises the combustion efficiency, tovirtually eliminate unburnt fuel, and oxidises the carbon monoxide to non-toxiccarbon dioxide, thus reducing air pollution.

The introduction of the air spray type burner to this type of combustion chamberalso greatly improves the preparation of fuel for combustion by aerating the over-rich pockets of fuel vapour close to the burner; this results in a large reduction ininitial carbon formation.

A high by-pass ratio engine will also reduce air pollution, since for a given thrustthe engine burns less fuel.

An Air Spray Fuel Nozzle.

Figure 5.6.



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5.4.4 REVERSE FLOW COMBUSTION CHAMBERReverse flow combustion chambers are used where the engine length is critical orwhere the thrust of the engine is not being produced by the exhaust of the primaryair. They are often found on APU’s, turboprop and turbo-shaft engines or theirderivatives such as the ALF 502 and LF507 engines used in the BAE 146 and RJaircraft.

By wrapping the combustion chamber around other components such as turbinesthe length of the engine can be significantly reduced. Losses in thrust do occurdue to the changes in airflow and direction of pressure forces. This is not importantin the types of engine where they are used as the majority of the thrust is derived

by other sources.They are often found on engines with compound compressors, which have acentrifugal stages as the last stage of compression.

Reverse Flow Combustion Chamber.Figure 5.8.



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5.6 COMBUSTION EFFICIENCYThe combustion efficiency of most gas turbine engines at sea-level take-offconditions is 100 per cent which reduces to 98 per cent at altitude cruiseconditions. The values vary as shown in because of the reducing air pressure,temperature and fuel/air ratio.

5.7 COMBUSTION STABILITY

Combustion stability means smooth burning and the ability of the flame to remainalight over a wide operating range.

For any particular type of combustion chamber there is both a rich and a weak limit

to the air/fuel ratio, beyond which the flame is extinguished. An extinction is mostlikely to occur in flight during a glide or dive with the engine idling, when there is ahigh airflow and only a small fuel flow, i.e. a very weak mixture strength.

The range of air/fuel ratio between the rich and weak limits is reduced with anincrease of air velocity, and if the air mass flow is increased beyond a certainvalue, flame extinction occurs. A typical stability loop is illustrated. The operatingrange defined by the stability loop must obviously cover the required air/fuel ratiosand mass flow of the combustion chamber.

The ignition process has weak and rich limits similar to those shown for stability.The ignition loop, however, lies within the stability loop, since it is more difficult to

establish combustion under ‘cold' conditions than to maintain normal burning.

Combustion Stability LimitsFigure 5.10.



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5.8 POLLUTION CONTROL

5.8.1 INTRODUCTIONPollution of the atmosphere by gas turbine engines falls into two categories; visible(ie. smoke) and invisible constituents (eg. oxides or nitrogen, unburnthydrocarbons, oxides of sulphur and carbon monoxide). The combination of thetraditional types of HP burner (eg. Duplex) with increasing compression ratios hasled to visible smoke during take-off and climb. The very strong public opinionagainst pollution of the atmosphere has forced engine manufacturers to developmethods of reducing smoke and other emissions.

5.8.2 SOURCES OF POLLUTION

Pollution occurs from incomplete combustion. When engines with high

compression ratios (ie. above 15:1) are fitted with the traditional type of atomisingburner, the high temperature, pressure and low turbulence within the combustionchamber prohibits adequate atomisation of the fuel when the engine is operatingat low altitude, thus causing the formation of carbon particles. This can bereduced to an acceptable level by improving the airflow inside the combustionchamber and by introducing burners that are not so susceptible to changes inpressure

5.9 EMISSIONS

The unwanted pollutants which are found in the exhaust gases are created withinthe combustion chamber. There are four main pollutants which are legislatively

controlled; unburnt hydrocarbons (unburnt fuel), smoke (carbon particles), carbonmonoxide and oxides of nitrogen. The principal conditions which affect theformation of pollutants are pressure, temperature and time.

In the fuel rich regions of the primary zone, the hydrocarbons are converted intocarbon monoxide and smoke. Fresh dilution air can be used to oxidise the carbonmonoxide and smoke into non-toxic carbon dioxide within the dilution zone.Unburnt hydrocarbons can also be reduced in this zone by continuing thecombustion process to ensure complete combustion.

Oxides of nitrogen are formed under the same conditions as those required for thesuppression of the other pollutants. Therefore it is desirable to cool the flame asquickly as possible and to reduce the time available for combustion. This conflictof conditions requires a compromise to be made, but continuing improvements incombustor design and performance has led to a substantially 'cleaner' combustionprocess.



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5.10 MATERIALSThe containing walls and internal parts of the combustion chamber must becapable of resisting the very high gas temperature in the primary zone. Inpractice, this is achieved by using the best heat resisting materials available, theuse of high heat resistant coatings and by cooling the inner wall of the flame tubeas an insulation from the flame.

The combustion chamber must also withstand corrosion due to the products of thecombustion, creep failure due to temperature gradients and fatigue due tovibrational stresses.

Methods of Cooling the Flame Tube.Figure 5.13.



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6 TURBINE SECTION

6.1 INTRODUCTION

The turbine has the task of providing the power to drive the compressor andaccessories and, in the case of engines which do not make use solely of a jet forpropulsion, of providing shaft power for a propeller or rotor. It does this by extractingenergy from the hot gases released from the combustion system and expandingthem to a lower pressure and temperature. High stresses are involved in thisprocess, and for efficient operation, the turbine blade tips may rotate at speeds over1,500 feet per second. The continuous flow of gas to which the turbine is exposedmay have an entry temperature between 850 and 1,700 deg.C. and may reach avelocity of over 2,500 feet per second in parts of the turbine.

A Triple Stage Turbine with a Single Shaft.Figure 6.1.



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To produce the driving torque, the turbine may consist of Several stages eachemploying one row of stationary nozzle guide vanes and one row of moving blades.The number of stages depends upon the relationship between the power requiredfrom the gas flow, the rotational speed at which it must be produced and the diameterof turbine permitted.

The number of shafts, and therefore turbines, varies with the type of engine., highcompression ratio engines usually have two shafts, driving high and low pressurecompressors. On high by pass ratio fan engines that feature an intermediatepressure system, another turbine may be interposed between the high and lowpressure turbines, thus forming triple-spool system. On some engines, driving torqueis derived from a free-power turbine. This method allows the turbine to run at itsoptimum speed because it is mechanically independent of other turbine and

compressor shafts.

A Multi Stage Turbine driving Two Shafts.

Figure 6.2.



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The mean blade speed of a turbine has considerable effect on the maximumefficiency possible for a given stage output. For a given output the gas velocities,deflections, and hence losses, are reduced in proportion to the square of highermean blade speeds. Stress in the turbine disc increases as the square of the speed,therefore to maintain the same stress level at higher speed the sectional thickness,hence the weight, must be increased disproportionately. For this reason, the finaldesign is a compromise between efficiency and weight. Engines operating at higherturbine inlet temperatures are thermally more efficient and have an improved powerto weight ratio. By-pass engines have a better propulsive efficiency and thus canhave a smaller turbine for a given thrust.

A Multi Stage Turbine Driving Three Shafts.

Figure 6.3.



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aerodynamic considerations, and to obtain optimum efficiency, compatible withcompressor and combustion design, the nozzle guide vanes and turbine blades areof a basic aerofoil shape. There are three types of turbine; impulse, reaction and acombination of the two known as impulse-reaction. In the impulse type the totalpressure drop across each stage occurs in the fixed nozzle guide vanes which,because of their convergent shape, increase the gas velocity whilst reducing thepressure. The gas is directed onto the turbine blades which experience an impulseforce caused by the impact of the gas on the blades. In the reaction type the fixednozzle guide vanes are designed to alter the gas flow direction without changing thepressure. The converging blade passages experience a reaction force resulting fromthe expansion and acceleration of the gas. Normally gas turbine engines do not usepure impulse or pure reaction turbine blades but the impulse-reaction combination.

The proportion of each principle incorporated in the design of a turbine is largelydependent on the type of engine in which the turbine is to operate, but in general it isabout 50 per cent impulse and 50 per cent reaction. Impulse-type turbines are usedfor cartridge and air starters.

6.2 ENERGY TRANSFER FROM GAS FLOW TO TURBINE

It will be seen that the turbine depends for its operation on the transfer of energybetween the combustion gases and the turbine. This transfer is never 100 per centbecause of thermodynamic and mechanical losses.

Comparison between a Pure Impulse Turbine and an Impulse Reaction Turbine.Figure 6.5.



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When the gas is expanded by the combustion process, it forces its way into thedischarge nozzles of the turbine where, because of their convergent shape, it isaccelerated to about the speed of sound which, at the gas temperature, is about2,500 feet per second. At the same time the gas flow is given a 'spin' or 'whirl' in thedirection of rotation of the turbine blades by the nozzle guide vanes. On impact withthe blades and during the subsequent reaction through the blades, energy isabsorbed, causing the turbine to rotate at high speed and so provide the power fordriving the turbine shaft and compressor.



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Twisted Contour of BladesFigure 6.6.

The torque or turning power applied to the turbine is governed by the rate of gas flowand the energy change of the gas between the inlet and the outlet of the turbineblades. The design of the turbine is such that the whirl will be removed from the gasstream so that the flow at exit from the turbine will be substantially 'straightened out'to give an axial flow into the exhaust system (Part 6). Excessive residual whirlreduces the efficiency of the exhaust system and also tends to produce jet pipevibration which has a detrimental effect on the exhaust cone supports and struts.

It will be seen that the nozzle guidevanes and blades of the turbine are'twisted', the blades having a staggerangle that is greater at the tip than atthe root. The reason for the twist is to

make the gas flow from thecombustion system do equal work atall positions along the length of theblade and to ensure that the flowenters the exhaust system with auniform axial velocity. This results incertain changes in velocity, pressureand temperature occurring through theturbine.

The 'degree of reaction' varies from

root to tip, being least at the root andhighest at the tip, with the meansection having the chosen value ofabout 50 per cent.

Gas Flow Pattern Through a Nozzle and Turbine.

Figure 6.7.



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The losses which prevent the turbine from being 100 per cent efficient are due to anumber of reasons. A typical uncooled three-stage turbine would suffer a 3.5 percent loss because of aerodynamic losses in the turbine blades. A further 4.5 per centloss would be incurred by aerodynamic losses in the nozzle guide vanes, gasleakage over the turbine blade tips and exhaust system losses; these losses are ofapproximately equal proportions. The total losses result in an overall efficiency ofapproximately 92 per cent.

6.3 CONSTRUCTION

The basic components of the turbine are the combustion discharge nozzles, thenozzle guide vanes, the turbine discs and the turbine blades. The rotating assembly

is carried on bearings mounted in the turbine casing and the turbine shaft may becommon to the compressor shaft or connected to it by a self-aligning coupling.

6.3.1 NOZZLE GUIDE VANES

The nozzle guide vanes are of an aerofoil shape with the passage between adjacentvanes forming a convergent duct. The vanes are located in the turbine casing in amanner that allows for expansion.The nozzle guide vanes are usually of hollow form and may be cooled by passingcompressor delivery air through them to reduce the effects of high thermalstresses and gas loads.

Typical Nozzle Guide Vane Construction.Figure 6.8.



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6.3.2 TURBINE DISCSTurbine discs are usually manufactured from a machined forging with an integralshaft or with a flange onto which the shaft may be bolted. The disc also has, aroundits perimeter, provision for the attachment of the turbine blades.

To limit the effect of heat conduction from the turbine blades to the disc a flow ofcooling air is passed across both sides of each disc.

6.3.3 TURBINE BLADES

The turbine blades are of an aerofoil shape, designed to provide passages betweenadjacent blades that give a steady acceleration of the flow up to the 'throat', where

the area is smallest and the velocity reaches that required at exit to produce therequired degree of reaction.

The actual area of each blade cross-section is fixed by the permitted stress in thematerial used and by the size of any holes which may be required for coolingpurposes. High efficiency demands thin trailing edges to the sections, but acompromise has to be made so as to prevent the blades cracking due to the

temperature changes during engine operation.

The method of attaching the blades to the turbine disc is of considerable importance,since the stress in the disc around the fixing or in the blade root has an importantbearing on the limiting rim speed. The blades on the early Whittle engine wereattached by the de Laval bulb root fixing, but this design was soon superseded by the'fir-tree' fixing that is now used in the majority of gas turbine engines. This type offixing involves very accurate machining to ensure that the loading is shared by all theserration’s. The blade is free in the serration’s when the turbine is stationary and isstiffened in the root by centrifugal loading when the turbine is rotating. Variousmethods of blade attachment are shown; however, the B.M.W. hollow blade and the

de Laval bulb root types are not now generally used on gas turbine engines.

Methods of Turbine Blade Attachment.Figure 6.9.



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A gap exists between the blade tips and casing, which varies in size due to thedifferent rates of expansion and contraction. To reduce the loss of efficiency throughgas leakage across the blade tips, a shroud is often fitted. This is made up by asmall segment at the tip of each blade which forms a peripheral ring around the bladetips. An abradable lining in the casing may also be used to reduce gas leakage.

Active Clearance Control (A.C.C.) is a more effective method of maintaining minimumtip clearance throughout the flight cycle. Air from the compressor is used to cool theturbine casing and when used with shroudless turbine blades, enables highertemperatures and speeds to be used.

Active Tip Clearance Control.Figure 6.10.



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6.3.4 DUAL ALLOY DISCS

Very high stresses are imposed on theblade root fixing of high work rate turbines,which make conventional methods of bladeattachment impractical. A dual alloy disc,or 'blisk' as shown in fig. 6.11., has a ringof cast turbine blades bonded to the disc.This type of turbine is suitable for smallhigh Power helicopter engines.

6.4 COMPRESSOR-TURBINE MATCHING

The flow characteristics of the turbine must be very carefully matched with those ofthe compressor to obtain the maximum efficiency and performance of the engine. If,for example, the nozzle guide vanes allowed too low a maximum flow, then a backpressure would build up causing the compressor to surge; too high a flow wouldcause the compressor to choke. In either condition a loss of efficiency would veryrapidly occur.

6.5 MATERIALS

Among the obstacles in the way of using higher turbine entry temperatures havealways been the effects of these temperatures on the nozzle guide vanes and turbineblades. The high speed of rotation which imparts tensile stress to the turbine discand blades is also a limiting factor.

6.5.1 NOZZLE GUIDE VANES

Due to their static condition, the nozzle guide vanes do not endure the samerotational stresses as the turbine blades. Therefore, heat resistance is the propertymost required. Nickel alloys are used, although cooling is required to prevent

melting. Ceramic coatings can enhance the heat resisting properties and, for thesame set of conditions, reduce the amount of cooling air required, thus improvingengine efficiency.

6.5.2 TURBINE DISCS

A turbine disc has to rotate at high speed in a relatively cool environment and issubjected to large rotational stresses. The limiting factor which affects the useful disclife is its resistance to fatigue cracking.

Section Through a Dual Alloy Disc.Figure 6.11.



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In the past, turbine discs have been made in ferritic and austenitic steels but nickelbased alloys are currently used. Increasing the alloying elements in nickel extend thelife limits of a disc by increasing fatigue resistance. Alternatively, expensive powdermetallurgy discs, which offer an additional 10% in strength, allow faster rotationalspeeds to be achieved.

6.5.3 TURBINE BLADES

A brief mention of some of the points to be considered in connection with turbineblade design will give an idea of the importance of the correct choice of bladematerial. The blades, while glowing red-hot, must be strong-enough to carry thecentrifugal loads due to rotation at high speed. A small turbine blade weighing onlytwo ounces may exert a load of over two tons at top speed and it must withstand thehigh bending loads applied by the gas to produce the many thousands of turbinehorsepower necessary to drive the compressor. Turbine blades must also beresistant to fatigue and thermal shock, so that they will not fail under the influence ofhigh frequency fluctuations in the gas conditions, and they must also be resistant tocorrosion and oxidisation. In spite of all these demands, the blades must be made ina material that can be accurately formed and machined by current manufacturingmethods-

From the foregoing, it follows that for a particular blade material and an acceptablesafe life there is an associated maximum permissible turbine entry temperature and acorresponding maximum engine power. It is not surprising, therefore, thatmetallurgists and designers are constantly searching for better turbine blade

materials and improved methods of blade cooling.

Over a period of operational timethe turbine blades slowly grow inlength. This phenomenon isknown as 'creep' and there is afinite useful life limit before failureoccurs.

Effect of Heat on Creep at Fixed Load.Figure 6.12.



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The early materials used were high temperature steel forgings, but these were rapidlyreplaced by cast nickel base alloys which give better creep and fatigue properties.

Close examination of a conventional turbine blade reveals a myriad of crystals that liein all directions (equi-axed). Improved service life can be obtained by aligning thecrystals to form columns along the blade length, produced by a method known as'Directional Solidification'. A further advance of this technique is to make the bladeout of a single crystal. Each method extends the useful creep life of the blade and inthe case of the single crystal blade, the operating temperature can be substantiallyincreased.

A non-metal based turbine blade can be manufactured from reinforced ceramics.Their initial production application is likely to be for small high speed turbines whichhave very high turbine entry temperatures.

Effect of Load on Creep at Constant Temperature.

Figure 6.13.



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Various turbine Blade Crystal Structures.Figure 6.14.



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Comparison of Turbine Blade Life Properties.Figure 6.15.



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6.6 DYNAMIC BALANCING PRINCIPLES6.6.1 INTRODUCTION

We must all be familiar with the effects of unbalance in one form or another, butperhaps the most common effect is that arising from wheel unbalance in motor cars. At resonance conditions it causes wobble or bounce, the effects of which aretransmitted to the driver through the steering column. This effect may be so violentas to make the car unsafe or at least uncomfortable to ride in, and the continualvibratory movements set up, even outside the resonance range will increase the rateof wear on the various linkages and add to driver and passenger fatigue.

In order to increase passenger comfort, reduce wear and noise levels and also to

increase the life of the engine between overhauls, design effort is put into the variousaspects of minimising vibration in aero-engines. Design features are also included topermit correction of unbalance forces.

Efforts are made to design engine bearing housings and carcasses with suitablestiffness to avoid resonance in the engine running range. In addition, precisebalancing instructions are issued to control the rotating forces on the bearings whichcould:-

a) be transmitted to other parts of the engine or airframe structure.

b) lead to engine failure in extreme cases.

The loads on the bearings are of three main forms. These are:

a) thrust loads due to the engine doing work.

b) journal loads due to the dead weight of engine parts.

c) unbalance loads.



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6.6.2 CENTRIFUGAL FORCE

Centrifugal force acts on every particle which makes up the mass of the rotatingelement impelling each particle outwards and away from the axis, about which it isrotating, in a radial direction.

If the mass of the rotating element is EVENLY DISTRIBUTED about the axis ofrotation, the part is BALANCED and rotates WITHOUT VIBRATION. However, ifthere is a greater mass on one side of the rotor than the other, the centrifugal forceacting on this heavy side exceeds the centrifugal force on the light side and pulls theentire assembly in the direction of the heavy side.

Centrifugal Forces.Figure 6.16..

Eccentric Mass.Figure 6.17.



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The rotor has a heavy mass M on one side. The centrifugal force exerted by Mcauses the entire rotor to be pulled in the direction of force F.

6.6.3 CAUSES OF UNBALANCE

Unbalance may be caused by a variety of factors occurring singly or in combinationwith others. These factors include:-

a) Eccentricity

Eccentricity exists when the geometric centreline of a part or assembly does notcoincide with its axis of rotation. This may be as a result of locating features (eg.spigot location, bolt holes, splines, serration’s, couplings), being eccentric to thebearing location.

b) Variation in Wall Thickness

Variation in wall thickness may be as a result of eccentricity between an inner andouter diameter of a cylindrical type feature, or it may be as a result of a difference inthickness between a radial section of a disk type feature and the section diametricallyopposite.

Eccentricity.Figure 6.18.

Variation in Wall Thickness.Fi ure 6.19.



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c) Blade DistributionUnbalance can be caused by an unequal or unsymmetrical arrangement of a set ofblades, either by reference to their mass moments or their dead weights dependingon the size of the blades. This can be as a result of faulty weighting, inaccurate orillegible recording or assembly errors.

d) Unsymmetrical Features

These may be due to manufacturing processes, such as blow holes in castings ordesign features such as offset holes, locating dogs, slots, keyways, etc.

e) Distortion

This can be caused by stress relieving, eg. after welding, or by unequal thermalgrowth during running.

f) Fits and Clearances

Clearance between mating parts allows relative movement of the parts and aconsequent shift of the axis of rotation during running (or even during balancing).Joints incompletely assembled, eg. chamfers fouling radii, abutment faces not pulledtogether, may cause a ‘bent’ rotor or an unsuitable joint, which may cause a shiftduring running. It is important to prevent separate locating, or fixing, features frominfluencing each other eg. bolt holes, spigot locations, serration’s, etc. must begeometrically controlled to prevent ‘fighting’ between more than one feature. Seealso the section on tooling, adapters, drives, dummy rotors, etc.

Unsymmetrical FeaturesFigure 6.20.



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g) Swash

Swash is caused by out of squareness of abutment faces relative to the bearingdiameter, abutment faces not being parallel across the component, eg. spacers,adjusting washers, disks, etc. It is important that the bolted joints are tightened insequence and in increments according to the torquing instructions.

h) Miscellaneous

Foreign bodies inside assemblies, oil accumulation, carbon deposits, usually foundwhen check balancing after running.

6.6.4 OBJECTIVE OF BALANCING

The objective of balancing is to determine how the unbalanced mass of the rotormust be compensated for in order to keep the bearings free of centrifugal forceloading.

6.6.5 DEFINITION OF UNBALANCE

Unbalance can be defined as that condition which exists in a rotor when vibratoryforce or motion is imparted to its bearings as a result of centrifugal forces.Unbalance will, in general, be distributed throughout the rotor but can be reduced to:-

a) static unbalance

b) couple unbalance

c) dynamic unbalance

Swash.Figure 6.21.



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Couple Unbalance

This arises when two EQUAL unbalance masses are positioned at opposite ends of arotor and spaced at 180 from each other. If placed on knife-edges, the rotor wouldbe statically balanced. However, when the rotor is rotated, the out of balancemasses will cause a centrifugal force to act at each end and hence each end willvibrate independently as shown in figure 6.23.

Dynamic UnbalanceThis occurs when the unbalanced masses may be either unequal in size or

positioned at some angle other than 180 to each other, or even both of theseconditions. These unbalanced forces now cause the rotor to vibrate.

Couple Unbalance.Figure 6.23.



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6.6.6 FAN BALANCINGBefore we look at fan balancing we must first look at vibration analysis techniquesadopted on modern gas turbines and the reason for doing it. One of therequirements of an on-condition maintenance policy is that defects can be detectedsufficiently early to permit rectification before secondary damage occurs. Theanalysis of engine vibration signatures is becoming an increasingly important tool fordetecting early failure in mechanical components.

A vibration monitoring system begins with a sensor, which may be a velocitytransducer or a peizo electric accelerometer. They both convert the mechanicalvibration of the machine into an electrical signal proportional to the vibrations

produced and together with the associated electrical circuitry feed signals to eithercockpit mounted gauges warning systems or a separate vibration analyser.

Velocity TransducerThis device operates on the principle of a permanent magnet to move within a coil,inducing voltage. Because of the moving parts with all the inherent disadvantages ofwear, friction, etc. they have been superseded by the peizo electric principle.

Peizo Electric AccelerometerIn this device, vibrating forces are transmitted to a peizo electric disc the resultantdeformation of the disc produces an electrical charge. Accelerometers have a

greater frequency range than velocity transducers and their lack of moving partsmakes them a much more stable and reliable means of collecting the basic vibrationsignal.

Many different specifications for accelerometers and transducers are available andsome of the considerations which govern their choice are:-

(1) DYNAMIC RANGE. The amplitude range over which the device is required toperform.

(2) SENSITIVITY. The severity of the vibration liable to be encountered.

(3) FREQUENCY RESPONSE. The full operating frequency range required.

(4) TEMPERATURE RANGE. The upper and lower temperature extremities towhich the device will be subjected and also any heat soak conditions.



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Peizo Electric Transducer.Fi ure 6.24.



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Figure 6.24. shows a schematic diagram of a typical peizo electric accelerometer.The top nut is torque loaded to give the correct starting datum on the peizo crystal.When subjected to a force (caused by engine vibration) the piezo electric crystalproduces an electric charge on its opposite faces. The output is fed to a chargeamplifier, which produces the voltage required for the cockpit indicator or frequencyanalyser. Most modern transducers employ a synthetic piezo electric such as leadzirconate in preference to natural quartz crystal because of the higher sensitivity forthe same force. In many cases, however, the choice of transducer will be dictated bythe operating temperature. The maximum allowable temperature for transducers is

typically 260C so they have to be sited on fan casings or in the by-pass ducting.

Transducers may be fitted in more than one plane or more than one location. Theanalyser can then be used to select a ‘broadband’ or overall vibration measurement,

which will give a quick guide to the condition of the engine.

Vibration monitoring varies greatly from aircraft to aircraft. The operator’srequirements and the technology of the aircraft will dictate the equipment fitted.Large commercial aircraft will have fitted a flight deck indication of the vibration levelsof engine spools, N1, N2, N3. Their main function is to warn the crew of a malfunction,ie. shed blade. The sensitivity of the vibration sensors may not be good enough fordetailed condition monitoring or fan balancing. Extra vibration sensors are fitted toenable these functions to be carried. There are some modern aircraft, which willcarry as a permanent fixture, eg. equipment that can carry out all vibration analysisrequirement.



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7 EXHAUST

7.1 INTRODUCTION

Aero gas turbine engines have an exhaust system which passes the turbinedischarge gases to atmosphere at a velocity, and in the required direction, to providethe resultant thrust. The velocity and pressure of the exhaust gases create the thrustin the turbo-jet engine, but in the turbo-propeller engine only a small amount of thrustis contributed by the exhaust gases, because most of the energy has been absorbedby the turbine for driving the propeller. The design of the exhaust system therefore,exerts a considerable influence on the performance of the engine. The areas of the jet pipe and propelling or outlet nozzle affect the turbine entry temperature, the massairflow and the velocity and pressure of the exhaust jet.

The temperature of the gas entering the exhaust system is between 550 and 850deg.C. according to the type of engine and with the use of afterburning can be 1,500deg.C. or higher. Therefore, it is necessary to use materials and a form ofconstruction that will resist distortion and cracking, and prevent heat conduction tothe aircraft structure.

A basic exhaust system is shown in fig. 7.1. The use of a thrust reverser, noisesuppressor and a two position propelling nozzle entails a more complicated systemas shown in fig. 7.2. The low by-pass engine may also include a mixer unit toencourage a thorough mixing of the hot and cold gas streams.

A Basic Exhaust System.Figure 7.1.



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An Exhaust System with a Thrust Reverser and Variable area propelling nozzle.Figure 7.2.



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7.2 EXHAUST GAS FLOWGas from the engine turbine enters the exhaust system at velocities from 750 to1,200 feet per second but, because velocities of this order produce high frictionlosses, the speed of flow is decreased by diffusion. This is accomplished by havingan increasing passage area between the exhaust cone and the outer wall as shownin fig. 7.3. The cone also prevents the exhaust gases from flowing across the rearface of the turbine disc. It is usual to hold the velocity at the exhaust unit outlet to aMach number of about 0.5, i.e. approximately 950 feet per second. Additional lossesoccur due to the residual whirl velocity in the gas stream from the turbine. To reducethese losses, the turbine rear struts in the exhaust unit are designed to straighten outthe flow before the gases pass into the jet pipe.

The exhaust gases pass to atmosphere through the propelling nozzle, which is a

convergent duct, thus increasing the gas velocity. In a turbo-jet engine, the exitvelocity of the exhaust gases is subsonic at low thrust conditions only. During mostoperating conditions, the exit velocity reaches the speed of sound in relation to theexhaust gas temperature and the propelling nozzle is then said to be 'choked'; that is,no further increase in velocity can be obtained unless the temperature is increased. As the upstream total pressure is increased above the value at which the propellingnozzle becomes ‘choked', the static pressure of the gases at the exit increases aboveatmospheric pressure. This pressure difference across the propelling nozzle giveswhat is known as 'pressure thrust' and is effective over the nozzle exit area. This isadditional thrust to that obtained due to the momentum change of the gas stream.

Exhaust Cone DetailFigure 7.3.



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With the convergent type ofnozzle a wastage of energyoccurs, since the gases leavingthe exit do not expand rapidlyenough to immediately achieveoutside air pressure.Depending on the aircraft flightplan, some high pressure ratioengines can with advantageuse a convergent-divergentnozzle to recover some of thewasted energy This nozzle

utilises the pressure energy toobtain a further increase in gasvelocity and, consequently, anincrease in thrust.

From the illustration (fig. 7.4), it will be seen that the convergent section exit nowbecomes the throat, with the exit proper now being at the end of the flared divergent

section. When the gas enters the convergent section of the nozzle, the gas velocityincreases with a corresponding fall in static pressure. The gas velocity at the throatcorresponds to the local sonic velocity. As the gas leaves the restriction of the throatand flows into the divergent section, it progressively increases in velocity towards theexit. The reaction to this further increase in momentum is a pressure force acting onthe inner wall of the nozzle. A component of this force acting parallel to thelongitudinal axis of the nozzle produces the further increase in thrust.

The propelling nozzle size is extremely important and must be designed to obtain thecorrect balance of pressure, temperature and thrust. With a small nozzle thesevalues increase, but there is a possibility of the engine surging, whereas with a large

nozzle the values obtained are too low. A fixed area propelling nozzle is only efficient over a narrow range of engineoperating conditions. To increase this range, a variable area nozzle may be used(Fig. 7.2.). This type of nozzle is usually automatically controlled and is designed tomaintain the correct balance of pressure and temperature at all operating conditions.In practice, this system is seldom used as the performance gain is offset by theincrease in weight. However, with afterburning a fully variable area nozzle isnecessary.

The by-pass engine has two gas streams to eject to atmosphere, the cool by-passairflow and the hot turbine discharge gases.

Gas Flow Through a Convergent Divergent NozzleFigure 7.4.



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In a low by-pass ratio engine, the two flows are combined by a mixer unit (fig. 7.5.)which allows the by-pass air to flow into the turbine exhaust gas flow in a manner thatensures thorough mixing of the two streams.

Low By-pass Mixer

Figure 7.5.



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In high by-pass ratio engines, the two streams are usually exhausted separately.The hot and cold nozzles are co-axial and the area of each nozzle is designed toobtain maximum efficiency. However, an improvement can be made by combiningthe two gas flows within a common, or integrated, nozzle assembly. This partiallymixes the gas flows prior to its ejection to atmosphere. An example of both types ofhigh by-pass exhaust system is shown in fig. 7.6.

High By-pass Engine Exhaust Systems.Figure 7.6.



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7.3 CONSTRUCTION AND MATERIALS

The exhaust system must be capable of withstanding the high gas temperatures andis therefore manufactured from nickel or titanium. It is also necessary to prevent anyheat being transferred to the surrounding aircraft structure. This is achieved bypassing ventilating air around the jet pipe, or by lagging the section of the exhaustsystem with an insulating blanket. Each blanket has an inner layer of fibrousinsulating material contained by an outer skin of thin stainless steel, which is dimpledto increase its strength. in addition, acoustically absorbent materials are sometimesapplied to the exhaust system to reduce engine noise.When the gas temperature is very high (for example, when afterburning is employed),

the complete jet pipe is usually of double-wall construction with an annular spacebetween the two walls. The hot gases leaving the propelling nozzle induce, byejector action, a flow of air through the annular space of the engine nacelle. Thisflow of air cools the inner wall of the jet pipe and acts as an insulating blanket byreducing the transfer of heat from the inner to the outer wall.

The cone and streamline fairings in the exhaust unit are subjected to the pressure ofthe exhaust gases; therefore, to prevent any distortion, vent holes are provided toobtain a pressure balance.

The mixer unit used in low by-pass ratio engines consists of a number of chutesthrough which the by-pass air flows into the exhaust gases. A bonded honeycomb

structure is used for the integrated nozzle assembly of high by-pass ratio engines togive lightweight strength to this large component.

Due to the wide variations of temperature to which the exhaust system is subjected, itmust be mounted and have its sections joined together in such a manner as to allowfor expansion and contraction without distortion or damage.

An Insulation BlanketFigure 7.7



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Exhaust JetJet noise is an externally generated source, which radiates in a rearward direction. Itis caused by the mixing process of the high-speed exhaust gases with thesurrounding air. In the mixing regions, a severe gradient of velocity exists normal tothe jet and due to the viscosity of the air, this gradient produces vortices and shearforces which, in turn, produce quadrupole noise sources.

The noise produced by such a source will be proportional to p2V je8, where p is the air

density and V je is the jet efflux velocity.

Noise Production in Sub & Super Sonic Air Flows.Figure 7.9



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Turbine Noise

Noise from the turbine is made up from two sources:

a. White Noise. “White”, random or background noise is caused by the reaction ofeach blade to the passage of air over its surface. There will always be noise fromeddy shedding in the blade wake reacting back on the blade and causing randomfluctuations over the blade surface (this source of noise may be likened to thatproduced by opening the quarter-light window on a car). Random noise will alsobe caused by turbulence in the air stream, which is sensed by the blade as achange in incidence with corresponding lift fluctuations and hence noise.

b. Discrete Noise. Discrete noise is produced by the regular passage of rotatingblades through the wakes from the preceding stationary vanes. If the spacebetween vanes and blades is small, there is a cyclic interaction between pressure

field. This can be overcome to some extent by design, ie. increasing the space. An additional source of discrete tones is caused by the rotating stage sensingchanges of incidence and hence lift pressure, passing through the wakes of theupstream vanes.

Compressor and Fan NoiseCompressor noise whilst significant, was relatively small compared with the exhaustnoise generated by turbojet and low by-pass engines. However as fans have gotlarger and by-pass ratios have increased the noise generated by the fan andcompressor may well exceed that produced by the exhaust.

Typical Quadrupole Noise Sources.Figure 7.10.



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Resultant Engine NoiseNoise from an engine is the combination of noises produced by the compressor, theturbine and the nozzle. With the low by-pass engine, the exhaust noise level dropsas the velocity of the exhaust gases is reduced and the turbine noise level drops asLP turbine mass flows and velocities are relatively reduced; but LP compressor noisebecomes significant over a wider range of thrust. As the by-pass ratio is increased,the exhaust jet and turbine noise levels continue to drop and the LP compressor (fan)noise level continues to rise. This trend continues until the exhaust jet noise level isless than the turbine noise level and the fan noise reaches a level comparable withexhaust jet of a pure jet engine. There will be no such increase in the fan noise if asingle-stage fan without IGV’s is aerodynamically suitable; instead, a significantdecrease to a level comparable to the turbine noise will occur, as illustrated in thefigure 7.12. This is because the more powerful elements of discrete tone andbackground noise are obviated.

Noise SuppressionIt has been seen that the first step towards noise suppression is at the design stageof the rotating and static parts of the engine. Thereafter, further reduction in thenoise level emanating from a particular engine may be achieved by the incorporationof special materials and innovations during its construction. These additionalmethods of noise suppression are briefly described as:

a) Absorption by acoustic linings.

b) Turbine, compressor and fan noise alleviated by control of nozzle area and

shape.

Comparison of Noise Sources of Low and High By-pass Engines.

Figure 7.12.



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Nozzle Area and Shape Control

In a high by-pass ratio engine with a single-stage fan without inlet guide vanes, thepredominant sources governing the overall noise level are the fan and turbine. If thefan speed can be reduced without loss of thrust, then the engine noise level would bereduced. At conditions below maximum thrust, the multi-spool engine enables this tobe accomplished by using a variable area nozzle to mechanically reduce the area ofthe hot stream final nozzle. This causes the speed of the LP turbine and itsassociated compressor spool to be reduced, producing a corresponding reduction infan and turbine noise levels. However, the velocity of the hot stream will increase,producing a corresponding rise in exhaust jet noise. If the final nozzle area isreduced until the noise level of the fan, turbine and exhaust are of the same order,the optimum mean noise level for the engine will have been achieved. This normally

occurs when the area of the hot stream final nozzle is reduced by approx. 50%. Atthe optimum nozzle area, the noise radiated towards the ground can be furtherreduced by a change in the geometrical shape of the nozzle.

Variable Area NozzleFigure 7.14.



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7.4.1.1 Exhaust Jet Mixing

Figure 7.12. shows that the noise from the exhaust jet is the main contributor to thetotal noise generated by a low by-pass ratio turbo-fan. For a turbo-jet the noise fromthe exhaust is an even greater contributor to the whole. Fortunately it iscomparatively easy to reduce the noise by increasing the mixture rate of the exhaustgases with the atmosphere. This can be achieved by increasing the contact area ofthe atmosphere with the gas stream by incorporating a corrugated or lobe-typesuppresser in the propelling nozzle.

The addition of a corrugate nozzlegives the effect shown in figure 7.16.

In the corrugated nozzle, atmosphericair flows down the outside corrugationsand into the exhaust jet to promoterapid mixing. In the lobe-type nozzle,the exhaust gases are divided to flowthrough the lobes and a small centralnozzle. This forms a number ofseparate exhaust jets which rapidly mixwith the air entrained by thesuppresser lobes. Deep corrugationsor lobes give a greater noise reduction,

but the penalties incurred limit the sizeof the suppressers, eg. to achieve therequired nozzle area, the overalldiameter of the suppresser may haveto be so large that excessive dragresults.

A nozzle may be designed to give alarge reduction in noise level, but thiscould incur a considerable weightpenalty due to the additional

strengthening required. A compromisethat gives a noticeable reduction innoise level with the minimum sacrificeof engine thrust or increase in weightis, therefore, the designer’s aim.

Type of Noise Suppressor.Figure 7.15.



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7.4.1.2 Recent Developments in Fan noise suppression

Rolls-Royce and GE are presently developing modified Trent and CF6 engines,respectively, which aim to reduce noise by incorporating chevron/saw tooth profiles totrailing edges of the fan and exhaust ducts. The manufacturers are alsoimplementing extended areas of acoustic nacelle lining. In the case of the Trentproof of-concept study, the acoustic liner area is increased by 30 per cent to 95 sq ft.

Improved Mixing by Corrugated Nozzle.Figure 7.16.



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The Rolls-Royce programme, in conjunction with Boeing, is already commencing testflights of a modified Trent 800-powered B777-200ER as part of an overall effort tocomply with ICAO Stage IV and QC2 noise levels - a pressing requirement necessaryfor future operations out of London's Heathrow airport. To this end, Rolls-Royce'snew technology may be applicable to B747s either on a retrofit or new-build basis,and the team expects jet noise reductions of at least 3 EPNdB at ground level.

Moreover, the modified fan case is expected to confer fan-noise reductions of 1.2EPNdB and 7 EPNdB from inside the cabin - particularly regarding the frequencieswhich cause a "fan-buzz" signature. GE meanwhile, is also targeting future Airbusand Boeing aircraft operations with its modified CF6 engine. This has been staticallytested in the autumn of 2002, with modified ducts and a new nozzle centrebody, forapplicability to existing A300/A310s.

According to GE, a peak jet noise reduction of 3.5dB is anticipated, while perceivedreductions are in the order of ldB. GE intends to implement the new configurationinto all its new-build CF6 engines from 2003. Like Rolls-Royce, GE is also targetingits big-fan modifications in conjunction with Boeing to facilitate ICAO Stage IV/QC2compliant B747 operations in the near future.

These serrated ducts will improve flow mixing and reduce noise on the Trent 800.

Figure 7.17.



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g. The reverser must not operate until required to do so. It is necessary to ensurethat:

1. Accidental selection of reverse thrust is impossible.

2. No single failure in the operating system selects reverse thrust.

3. The thrust changing elements are biased away from the reverse thrustposition.

7.5.3 LAYOUT AND OPERATION OF TYPICAL THRUST REVERSING SYSTEMS

Clamshell door systemThe clamshell door system is a pneumatically operated system, as shown in detail infig. 7.19. Normal engine operation is not affected by the system, because the ducts

through which the exhaust gases are deflected remain closed by the doors untilreverse thrust is selected by the pilot.

On the selection of reverse thrust, the doors rotate to uncover the ducts and close thenormal gas stream exit. Cascade vanes then direct the gas stream in a forwarddirection so that the jet thrust opposes the aircraft motion.

The clamshell doors are operated by pneumatic rams through levers that give themaximum load to the doors in the forward thrust position; this ensures effectivesealing at the door edges, so preventing gas leakage. The door bearings andoperating linkage operate without lubrication at temperatures of up to 600 deg.C.

Bucket target systemThe bucket target system is hydraulically actuated and uses bucket-type doors toreverse the hot gas stream. The thrust reverser doors are actuated by means of aconventional pushrod system. A single hydraulic powered actuator is connected to adrive idler, actuating the doors through a pair of pushrods (one for each door).

The reverser doors are kept in through the drive idler. The hydraulic actuatorincorporates a mechanical lock in the stowed (actuator extended) position.

Clamshell Doors.Figure 7.19.



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In the forward thrust mode (stowed) the thrust reverser doors form the convergent-divergent final nozzle for the engine.

Cold stream reverser systemThe cold stream reverser system can be actuated by an air motor, the output ofwhich is converted to mechanical movement by a series of flexible drives, gearboxesand screwjacks, or by a system incorporating hydraulic rams.

When the engine is operating in forward thrust, the cold stream final nozzle is 'open'because the cascade vanes are internally covered by the blocker doors (flaps) andexternally by the movable (translating) cowl; the latter item also serves to reducedrag.

On selection of reverse thrust, the actuation system moves the translating cowlrearwards and at the same time folds the blocker doors to blank off the cold stream

final nozzle, thus diverting the airflow through the cascade vanes.

Bucket Type Thrust Reverser.Figure 7.20.



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7.5.3.1 Combination Reversers

Some engines are equipped with both cold and hot stream reversers, these have thesome benefits of both types as well as some of the disadvantages.

Cold Stream Reverser.Figure 7.21.

Hot and Cold Stream Reverser.Figure 7.22.



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7.5.4 SAFETY FEATURESReverse thrust systems will have some of the following safety features incorporated:

a. Reverse thrust cannot be selected until the engine throttle is brought backto idle.

b. A mechanical lock prevents doors moving from the forward thrust positionuntil reverse thrust is selected.

c. Acceleration in forward thrust can only be obtained when the reversethrust lever is de-selected and the doors are in the open position.

d. Acceleration in reverse thrust can only be obtained when the reverse

thrust lever is selected and the doors are in the closed position.

e. The aircraft has to be on the ground or very close to it before reversethrust selection is allowed (this does not apply to aircraft that use reversethrust as an airbrake in flight).

On the cold stream reverser/hot stream spoiler system, a mechanical interlockprevents reverse thrust being selected except when the throttle lever is at the idleposition. After selection, acceleration of the engine to give reverse thrust isprevented until the translating cowl has moved rearwards. When the cowl hasmoved into position, a mechanical feedback from the cowl screw-jack unlocks thethrottle control.

7.5.5 CFM 56 THRUST REVERSER FOR BOEING 737-300

The 737-300 is equipped with electrically controlled, hydraulically powered, fan onlythrust reversers. The thrust reversers are interchangeable between the two enginesexcept for the cascade basket assemblies and the strikers which deflect the Kruegerflaps when the fan cowl translates aft.

Boeing 737-300 Thrust Reverser in Stowed and Deployed Positions.Figure 7.23



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Reverser actuation is controlled by a control valve module, located on the forwardbulkhead of each air-conditioning bay. This module contains two control valves(isolation and direction) and a manually operated (pinnable) maintenance shut-offvalve. The control valves are operated by solenoids which are actuated by the thrustlever switches.

Operation of the Blocker Doors.Figure 7. 24.

Reverser Control Valve Module.Figure 7.25.



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The thrust reverser hydraulic system is only pressurised when thrust reverseractuation is required, or when required to resist motion from the stow commandedposition.

Application of hydraulic power to the reversers by operation of the reverse thrustlevers is prevented unless the aeroplane is within 10 feet of the ground (radioaltimeter 1 or 2), or is on the ground (right-hand main gear oleo compressed). Pullingan engine fire handle prevents the isolation valve from opening, or closes it if it isalready open. A high idle is maintained for 4 seconds after activation of the weighton wheels switch in order to improve engine spool-up time in reverse.

Each thrust reverser is powered by a separate hydraulic system, with a standbysystem available as an alternate source with a reduced deployment rate.

An automatic restow system activates an actuator stow force anytime the reverser issensed to be out of the stowed position during forward thrust operation.

Location of T/R Actuators and Synchronisation System.

Figure 7.26.



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Thrust Reverser Schematic.Figure 7.27.



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A throttle interlock system restricts application of engine thrust when the reverser isnot in its commanded position and automatically reduces engine thrust ifuncommanded reverser translation occurs.

Amber lights on the centre panel identify when the reversers are in the unlockedposition.

A "fault light" for each reverser is located in the Engine Module on the aft overheadpanel. When this fault light is illuminated, the Master Caution is triggered after 12seconds to indicate that a subsequent failure in the reverser system may causeuncommanded reverser motion.



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Thrust Reverser Controls.Figure 7.28.

A

B



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8 BEARINGS, SEALS AND GEARBOXES

8.1 BEARINGS8.1.1 INTRODUCTION

A bearing is any surface that supports or is supported by another surface. Bearingsare designed to produce a minimum of friction and a maximum of wear resistance.

Bearings must reduce the friction of moving parts and also take thrust loads or acombination of thrust and radial loads. Those which are designed primarily forthrust loads are called thrust bearings. The ball bearings are used to provide thethrust bearing as they can take both thrust and radial loads, and roller bearings areused to support the shaft whilst allowing axial movement. They are sometimescalled expansion bearings.

8.1.2 BALL BEARINGS

A ball bearing consists of an inner race, an outer race and one or more sets of balls,and a ball retainer or cage. The purpose of the retainer or cage is to prevent theballs touching one another. Ball bearings are used for radial and thrust loads; a ballbearing specially designed for thrust loads would have very deep grooves in theraces or be of the angular bearing type, these must always be fitted the correct wayround!

8.1.3 ROLLER BEARINGS

These bearings are manufactured in various shapes and sizes and will withstand

greater radial loads than ball bearings because of greater contact area. They allowaxial movement of the shaft, this is very useful in a gas turbine due to expansion ofthe engine due to the heat it produces.

8.1.4 OTHER TYPES OF BEARINGS

It is rare to find taper roller or needle bearings used in gas turbine engines, howeversome APU’s use plain bearings to support the turbine end of the main shaft.



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Plain Roller BearingFigure 8.1.

Examples of Bearing Types.Figure 8.2.



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8.2 BEARING CHAMBER OR SUMPOne or more bearings are contained within a bearing chamber or sump. Thechamber is sealed to prevent oil escaping into the engine and to prevent excessiveair getting into the oil.

8.2.1 LUBRICATION

The bearing chamber will have an oil feed which is sprayed on to the bearing tolubricate and cool it.

On some engines, to minimise the effect of the dynamic loads transmitted from therotating assemblies to the bearing housings, a ‘squeeze film' type of bearing is

used. They have a small clearance between the outer race of the bearing andhousing with the clearance being filled with pressurised oil. The oil film dampensthe radial motion of the rotating assembly and the dynamic loads transmitted to thebearing housing thus reducing the vibration level of the engine and the possibility ofdamage by fatigue.

The oil will return to the oil system from the bottom of the bearing chamber, eitherby gravity or by suction from a scavenge pump.

8.2.2SEALING

Bearing chambers are usually sealed using air. The internal cooling air within theengine provides the air. Typical seals used are labyrinth, screw back and carbon

types. . All of these seals need a differential pressure between inside and outsidethe bearing housing . Where pressure is available it is used, if the differential is toolow, it can be boosted by suction from a scavenge pump. Carbon seals require theoil to be in contact with them to provide cooling for the seal.

To prevent excess pressure building up within the bearing chamber, it is usuallyvented. This vent on some engines is taken to the oil tank to ensure that the wholesystem is working against the same pressure, or it goes to the oil pressure regulatorto ensure that there is a constant pressure drop across the spray jets in the bearinghousings.

8.2.2.1 Labyrinth Seals

Labyrinth seals are constructed of metal non-rotating lands, which are secured tovarious parts of the engine case and a series of cylindrical rotating knife-edge stepsthat mate with the lands. With this type of seal, there are no contacting parts. Aprecise clearance is designed into the seals to control the pressure, as thecompressor air passes over the cascade of knife-edges, the pressure is reduced.



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8.2.4 CARBON SEAL Another method of air sealing is achieved by using a carbon seal arrangement.They are used on the rotating assembly of a gas turbine and protection of enginedrive components in accessory gearboxes.

Carbon seals are manufactured of a mixture of carbon and graphite powder,bonded together with a viscous substance, such as coal tar. The carbon seal isfixed and held against the rotating seal by springs. Both the rotating seal and thecarbon seals are machine ground and precision lapped to a micro finish.

8.2.5 SPRING RING SEAL

This type of seal would normally be used around a main bearing assembly withinthe engine. It may be used in conjunction with a labyrinth or screw back type ofseal.

The ring seal is similar to a large stepped pistonring; it is located on a rotating shaft. When theshaft is stationary, the seal clamps tightly to theshaft. As the shaft rotates, the spring ring canexpand slightly, under centrifugal force, when itthen forms an effective seal with the adjacentstationary housing.

Carbon Seal.Figure 8.5.

Ring Seal

Figure 8.6.



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8.3 ACCESSORY DRIVE GEARBOXES

8.3.1 INTRODUCTIONGearboxes provide the power for aircraft hydraulic, pneumatic and electricalsystems in addition to providing various pumps and control systems for efficientengine operation. The high level of dependence upon these units requires anextremely reliable drive system.

The drive for the gearbox is typically taken from a rotating engine shaft usually theHP shaft, via an internal gearbox, to an external gearbox that provides a mount forthe accessories and distributes the appropriate geared drive to each accessory. Astarter may also be fitted to provide an input torque to the engine. An accessorydrive system on a high by-pass engine takes between 400 and 500 horsepower

from the engine.8.3.2 INTERNAL GEARBOX

The location of the internal gearbox within the core of an engine is dictated by thedifficulties of bringing a driveshaft radially outwards and the space available withinthe engine core.

Mechanical Arrangements of Accessory Drive Gearboxes.Figure 8.8.



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Thermal fatigue and a reduction in engine performance, due to the radial driveshaftdisturbing the gasfiow, create greater problems within the turbine area than thecompressor area. For any given engine, which incorporates an axial-flowcompressor, the turbine area is smaller than that containing the compressor andtherefore makes it physically easier to mount the gearbox within the compressorsection. Centrifugal compressor engines can have limited available space, so theinternal gearbox may be located within a static nose cone or, in the case of a turbo-propeller engine, behind the propeller reduction gear as shown in fig.8.8.

On multi-shaft engines, the choice of whichcompressor shaft is used to drive theinternal gearbox is primarily dependentupon the ease of engine starting. This is

achieved by rotating the compressor shaft,usually via an input torque from the externalgearbox. In practice the high pressuresystem is invariably rotated in order togenerate an airflow through the engine andthe high pressure compressor shaft istherefore coupled to the internal gearbox.

To minimise unwanted movement betweenthe compressor shaft bevel gear and radialdriveshaft bevel gear, caused by axial

movement of the compressor shaft, thedrive is taken by one of three basicmethods (fig. 8.9.). The least number ofcomponents is used when the compressorshaft bevel gear is mounted as close to thecompressor shaft location bearing aspossible, but a small amount of movementhas to be accommodated within themeshing of the bevel gears. Alternatively,the compressor shaft bevel gear may bemounted on a stub shaft that has its own

location bearing. The stub shaft is splinedonto the compressor shaft that allows axialmovement without affecting the bevel gearmesh. A more complex system utilises anidler gear that meshes with the compressorshaft via straight spur gears,accommodating the axial movement, anddrives the radial driveshaft via a bevel geararrangement. The latter method was widelyemployed on early engines to overcomegear engagement difficulties at high speed.Types of Internal Gearbox

Figure 8.9.



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To spread the load of driving accessory units, some engines take a second drivefrom the slower rotating low pressure shaft to a second external gearbox (fig.8.8.).This also has the advantage of locating the accessory units in two groups, thusovercoming the possibility of limited external space on the engine. When thismethod is used, an attempt is made to group the accessory units specific to theengine onto the high pressure system, since that is the first shaft to rotate, and theaircraft accessory units are driven by the low pressure system. A typical internalgearbox showing how both drives are taken is shown in fig.8.10. This method mayalso be used to drive speed sensors and governors for the low pressure shaft.

An Internal Gearbox With an LP and HP Output.Figure 8.10.



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8.3.3 RADIAL DRIVESHAFTThe purpose of a radial driveshaft is to transmit the drive from the internal gearboxto an accessory unit or the external gearbox. It also serves to transmit the hightorque from the starter to rotate the high pressure system for engine startingpurposes. The driveshaft may be direct drive or via an intermediate gearbox.

To minimise the effect of the driveshaft passing through the compressor duct anddisrupting the airflow, it is housed within the compressor support structure. On by-pass engines, the driveshaft is either housed in the outlet guide vanes or in a hollowstreamlined radial fairing across the low pressure compressor duct.

To reduce airflow disruption it is desirable to have the smallest driveshaft diameter

as possible. The smaller the diameter, the faster the shaft must rotate to providethe same power. However, this raises the internal stress and gives greater dynamicproblems, which result in vibration. A long radial driveshaft usually requires a rollerbearing situated halfway along its length to give smooth running. This allows arotational speed of approximately 25,000 r.p.m. to be achieved with a shaft diameterof less than 1.5 inch without encountering serious vibration problems.

8.3.4 DIRECT DRIVE

In some early engines, a radial driveshaft was used to drive each, or in someinstances a pair, of accessory units. Although this allowed each accessory unit tobe located in any desirable location around the engine and decreased the power

transmitted through individual gears, it necessitated a large internal gearbox. Additionally, numerous radial driveshafts had to be incorporated within the design.This led to an excessive amount of time required for disassembly and assembly ofthe engine for maintenance purposes.

In some instances the direct drive method may be used in conjunction with theexternal gearbox system when it is impractical to take a drive from a particular areaof the engine to the external gearbox. For example, figure8.8. shows a turbo-propeller engine which requires accessories specific to the propeller reduction drive,but has the external gearbox located away from this area to receive the drive fromthe compressor shaft.

8.3.5 GEAR TRAIN DRIVE

When space permits, the drive may be taken to the external gearbox via a gear train(fig.8.8). This involves the use of spur gears, sometimes incorporating a centrifugalbreather. However, it is rare to find this type of drive system in current use.

8.3.6 INTERMEDIATE GEARBOX

Intermediate gearboxes are employed when it is not possible to directly align theradial driveshaft with the external gearbox. To overcome this problem anintermediate gearbox is mounted on the high pressure compressor case and re-

directs the drive, through bevel gears, to the external gearbox. An example of thislayout is shown in fig.8.8.



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8.3.7 EXTERNAL GEARBOX

The external gearbox contains the drives for the accessories, the drive from thestarter and provides a mounting face for each accessory unit. Provision is alsomade for hand turning the engine, via the gearbox, for maintenance purposes.Fig.8.11. shows the accessory units that are typically found on an external gearbox.

An External Gearbox.Figure 8.11.



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An External Gearbox with an Auxiliary Gearbox Drive.Figure 8.12.



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8.3.9CONSTRUCTION AND MATERIALS

Gears

The spur gears of the external or auxiliary gearbox gear train (figs.8.11. and 8.12.)are mounted between bearings supported by the front and rear casings which arebolted together. They transmit the drive to each accessory unit, which is normallybetween 5000 and 6000 r.p.m. for the accessory units and approximately 20,000r.p.m. for the centrifugal breather.

All gear meshes are designed with 'hunting tooth' ratios which ensure that eachtooth of a gear does not engage between the same set of opposing teeth on each

revolution. This spreads any wear evenly across all teeth.Spiral (helical) bevel gears are used for the connection of shafts whose axes are atan angle to one another but in the same plane. The majority of gears within a geartrain are of the straight spur gear type, those with the widest face carry the greatestloads. For smoother running, helical gears are used but the resultant end thrustcaused by this gear tooth pattern must be catered for within the mounting of thegear.

Gearbox sealing

Sealing of the accessory drive system is primarily concerned with preventing oilloss. The internal gearbox has labyrinth seals where the static casing mates with

the rotating compressor shaft. For some of the accessories mounted on theexternal gearbox, an air blown pressurised labyrinth seal is employed. Thisprevents oil from the gearbox entering the accessory unit and also preventscontamination of the gearbox, and hence engine, in the event of an accessoryfailure. The use of an air blown seal results in a gearbox pressure of about 3 lbs.per sq. in. above atmospheric pressure. To supplement a labyrinth seal, an 'oilthrower ring' may be used. This involves the leakage oil running down the drivingshaft and being flung outwards by a flange on the rotating shaft. The oil is thencollected and returned to the gearbox.

Materials

To reduce weight, the lightest materials possible are used. The internal gearboxcasing is cast from aluminium but the low environmental temperatures that anexternal gearbox is subjected to allows the use of magnesium castings which arelighter still, The gears are manufactured from non-corrosion resistant steels forstrength and toughness. They are case hardened to give a very hard wear resistantskin and feature accurately ground teeth for smooth gear meshing.



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9 LUBRICANTS AND FUEL

9.1 GAS TURBINE FUEL PROPERTIES AND SPECIFICATION

Introduction

In the earliest days of the gas turbine engine, kerosene was regarded as the mostsuitable fuel. It commended itself on the grounds of availability, cost, calorific value,burning characteristics and low fire hazard.

Other types of petroleum fuels are not suitable for use in gas turbines because ofthe wide range of temperature and pressure over which combustion must occur andthe necessity of keeping the weight and volume to a minimum.

General Requirements

A gas turbine fuel should have the following qualities:

a) Ease of flow under all operating conditions.

b) Quick starting of the engine.

c) Complete combustion under all conditions.

d) A high calorific value.

e) Non-corrosive.

f) The by-products of combustion should have no harmful effect on the flametubes, turbine blades, etc.

g) Minimum fire hazards.

h) Provide lubrication of the moving parts of the fuel system.

i) The by-products of combustion should have minimal harmful effect on theenvironment

9.2 FRACTIONAL DISTILLATION

This process is carried out in a fractionating column, which has a series of trays asshown in the figure. The effect of the superheated steam on the heated crudepetroleum is to cause the lighter fractions to rise up the column. When rising, the

vapour cools and a certain amount condenses on each tray until the tray is full ofliquid to the overflow. Thus, each tray is a little cooler than the one below it, andtherefore, lighter and lighter fractions will be present on each tray, as the vapourspass up the column. The temperature is controlled at the bottom of the column bythe temperature of the crude oil, and at the top of the column by taking a certainamount of the product as it leaves, condensing it and pumping it back into the top ofthe column. This is known as the reflux.

A certain amount of material will condense, which has a lower boiling point than thebulk of the liquid on a particular tray. To enable separation of these fractions, theliquid from a selected tray is drawn into a smaller auxiliary column, called a ‘side-stripper’. Here it is treated with steam that causes the lightest fractions to vaporise

and pass along with the steam into the main column.



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Fractioning Tower.Figure 9.1.



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The use of these side-strippers enables kerosene and gas oil to be obtained directfrom the plant. Lubricating oil distillate, if such is present, can usually be drawndirect from a tray without the use of a side-stripper, while gasoline leaves the top ofthe column as a vapour and must be cooled to condense it to liquid gasoline.

9.3 PROPERTIES

9.3.1 EASE OF FLOW

The ease of flow of a fuel is mainly a question of viscosity, but impurities such asice, dust, wax, etc., may cause blockages in filters and in the fuel system generally.

Most liquid petroleum fuels dissolve small quantities of water and if the temperatureof the fuel is reduced enough, water or ice crystals are deposited from the fuel. Adequate filtration is therefore necessary in the fuel system. The filters may have to

be heated, or a fuel de-icing system fitted, to prevent ice crystals blocking the filters.Solids may also be deposited from the fuel itself due to the solidification of waxes orother high molecular weight hydrocarbons. Distillates heavier than kerosene, suchas gas oil, generally have a pour point temperature too high for use in aircraftoperating in low temperatures. If these fuels were to be used, some form of heatingin the aircraft’s tanks and fuel system would be necessary. Such heating wouldobviously be an unreasonable complication.

9.3.2 EASE OF STARTING

The speed and ease of starting of gas turbines depends on the ease of ignition ofan atomised spray of fuel. This ease of ignition depends on the quality of the fuel in

two ways:

a) The volatility of the fuel at starting temperatures.

b) The degree of atomisation, which depends on the viscosity of the fuel as wellas the design of the atomiser.

The viscosity of fuel is important because of its effect on the pattern of the liquidspray from the burner orifice and because it has an important effect on the startingprocess. Since the engine should be capable of starting readily under all conditionsof service, the atomised spray of fuel must be readily ignitable at low temperatures.Ease of starting also depends on volatility, but in practice the viscosity is found to be

the more critical requirement. In general, the lower the viscosity and the higher thevolatility, the easier it is to achieve efficient atomisation.

9.3.3 COMPLETE COMBUSTION

The exact proportion of air to fuel required for complete combustion is called thetheoretical mixture and is expressed by weight. There are only small differences inignition limits for hydrocarbons, the rich limit in fuels of the kerosene range being5:1 air/fuel ratio by weight and the weak limit about 25:1 by weight.

Flammable air/fuel ratios each have a characteristic rate of travel for the flamewhich depends on the temperature, pressure and the shape of the combustionchamber. Flame speeds of hydrocarbon fuels are very low and range from 0.3 –

0.6 m/sec. These low values necessitate the provision of a region of low air velocitywithin the flame tube, in which a stable flame and continuous burning are ensured.



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Flame temperature does not appear to be directly influenced by the type of fuel,except in a secondary manner as a result of carbon formation, or of pooratomisation resulting from a localised over-rich mixture. The maximum flame

temperature for hydrocarbon fuels is roughly 2,000C. This temperature occurs at amixture strength slightly richer than the theoretical, owing to dissociation of themolecular products of combustion, which occurs at the theoretical mixture.

Dissociation occurs above about 1,400C and reduces the energy available fortemperature rise.

The problem of the flame becoming extinguished in flight is not perfectlyunderstood, but it appears that the type of fuel is of relatively minor importance.However, wide cut gasoline’s are more resistant to extinction than kerosene andengines are easier to relight using wide cut fuel. This is due to the higher vapour

pressure of these fuels.

9.3.4 CALORIFIC VALUE

The calorific value is a measure of the heat potential of a fuel. It is of greatimportance in the choice of fuel, because the primary purpose of the combustionsystem is to provide the maximum amount of heat with the minimum expenditure offuel. The calorific value of liquid fuels is usually expressed in megajoules (MJ) perlitre. When considering calorific value, it should be noted that there are two valueswhich can be quoted for every fuel, the gross value and the net value. The grossvalue includes the latent heat of vaporisation and the net value excludes it. The netvalue is the quantity generally used. The calorific value of petroleum fuels is related

to their specific gravity. With increasing specific gravity (heavier fuels) there is anincrease in calorific value per litre but a reduction in calorific value per kilogram.Thus, for a given volume of fuel, kerosene gives an increased aircraft range whencompared with gasoline, but weighs more. If the limiting factor is the volume of thefuel tank capacity, a high calorific value by volume is the more important.

9.3.5 CORROSIVE PROPERTIES

The tendency of a turbine fuel to corrode the aircraft’s fuel system depends on twofactors:-

a) Water.

b) Other corrosive substances, notably sulphur compounds.

The water which causes corrosion is dissolved water. It leads to corrosion of thefuel system, which is particularly important with regard to the sticking of slidingparts, especially those with small clearances and only small or occasionalmovement.

Corrosion can also be caused by secondary effects, such as biological corrosioncaused by plant spores, which are not killed off by the cracking process. Keroseneand diesel suffer from this form of contamination.



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9.3.6 EFFECTS OF BY-PRODUCTS OF COMBUSTION

Carbon deposition in the combustion system indicates imperfect combustion andmay lead to:-

a) A lowering of the surface temperature on which it is deposited, resulting inbuckled flame tubes because of the thermal stresses set up by thetemperature differences.

b) Damage to turbine blades caused by lumps of carbon breaking off and strikingthem.

c) Disruption of airflow through the turbine, creating turbulence, back-pressureand possible choking of the turbine, resulting in loss of efficiency.

It appears that carbon deposition depends on the design of the combustionchamber and the aromatic content of the fuel. (Aromatics are a series ofhydrocarbons based on the benzene ring). The higher the aromatic content, thegreater the carbon deposits.

Sulphur will affect the turbine. Every effort is made to keep the sulphur content aslow as possible in aviation turbine fuels. Unfortunately, removal of the sulphurinvolves increased refining costs and decreased supplies and so some sulphur istherefore permitted.

9.3.7 FIRE HAZARDS

There are three main sources of fire hazard; these arise from:-

a) Fuel spillage with subsequent ignition of the vapour from a spark, etc.

b) Fuel spillage on to a hot surface causing self-ignition.

c) The existence of inflammable or explosive mixtures in the aircraft fuel tanks.

The first hazard depends on the volatility of the fuel. The lower the flash point, thegreater the chance of fire through this cause. It is more difficult to ignite kerosenethan to ignite gasoline or wide cut fuel in this way.

The second hazard depends on the spontaneous ignition temperature of the fuel. Inthis respect, gasoline has a slightly higher spontaneous ignition temperature than

kerosene, but if a fire does occur, the rate of spread is much slower in keroseneowing to its lower volatility.

The third hazard depends upon the temperature and pressure in the tank and thevolatility of the fuel. For any fuel there are definite temperature limits within which aflammable fuel vapour/air mixture will exist. If the temperature falls below the lowerlimit, the mixture will be too weak to burn, while if the temperature rises above theupper limit, the mixture is too rich to burn. At ground level the comparativetemperature limits of flammability for gasoline and kerosene is as follows:

a) Gasoline. Upper limit -10C. Lower limit -42C.

b) Kerosene. Upper limit +90C. Lower limit +43C.



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At higher altitudes the temperatures are somewhat lower. This informationindicates that explosive conditions in the vapour space will occur with the lowvolatility turbine fuel under extremely hot weather conditions and with gasolineunder extremely low temperature conditions.

9.3.8 VAPOUR PRESSURE

The vapour pressure of a liquid is a measure of its tendency to evaporate. Thesaturated vapour pressure (SVP) of a liquid (ie. the pressure exerted by vapour incontact with the surface of the liquid) increases with increasing temperature. Whenthe SVP equals the pressure acting on the surface of the liquid, the liquid boils.Thus, the boiling point of a liquid depends on a combination of SVP, the pressureacting on its surface and its temperature.

9.3.9 FUEL BOILING AND EVAPORATION LOSSES At high rates of climb, fuel boiling and evaporation is a problem which is not easilyovercome. A low rate of climb permits the fuel in the tanks to cool and thus reduceits vapour pressure as the atmospheric pressure falls off. However, the rate ofclimb of many aircraft is so high that the fuel retains its ground temperatures, so thaton reaching a certain altitude the fuel begins to boil. In practice this boiling hasproved to be so violent that the loss is not confined to vapour alone. Layers ofbubbles form and are swept through the tank vents with the vapour stream. Thisloss is analogous to a saucepan boiling over and is sometimes referred to asslugging.

The amount of fuel lost from evaporation depends on several factors:a) Vapour pressure of the fuel.

b) Fuel temperature on take-off.

c) Rate of climb.

d) Final altitude of the aircraft.

Fuel losses as high as 20% of the tank contents have been recorded through boilingand evaporation.

9.3.10 METHODS OF REDUCING OR ELIMINATING FUEL LOSSES

Possible methods of reducing or eliminating losses by evaporation are:

a. Reduction of the rate of climb.

b. Ground cooling of the fuel.

c. Flight cooling of the fuel.

d. Recovery of liquid fuel and vapour in flight.

e. Re-design of the fuel tank vent system.

f. Pressurisation of the fuel tanks.



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Reduct ion of the Rate of Climb

Reducing the rate of climb imposes an unacceptable restriction on the aircraft anddoes not solve the problem of evaporation loss. This method is, therefore, notused.

Ground Cooling of the FuelThis is not considered a practical solution, but in hot climates every effort should bemade to shade refuelling vehicles and the tanks of parked aircraft.

Flight Cooling of the FuelThe use of a heat exchanger, through which the fuel is circulated to reduce thetemperature sufficiently to prevent boiling, is possible. High rates of climb, however,

would not allow enough time to cool the fuel without the aid of heavy or bulkyequipment. At a high true airspeed speeds TAS, the rise in airframe temperaturedue to skin friction increases the difficulty of using this method. On small high-speed aircraft the weight and bulk of the coolers becomes prohibitive.

Recovery of L iquid Fuel in FlightThis method would probably entail bulky equipment and therefore is unacceptable. Another method would be to convey the vapour to the engines and burn it toproduce thrust, but the complications of so doing would entail severe problems.

Redesign of the Fuel Tank Vent System

The loss of liquid fuel could be largely eliminated by redesigning the vents, but theevaporative losses would remain. However, improved venting systems may wellprovide a more complete solution to the problem.

Pressurisation of the Fuel TanksThere are two ways in which fuel tanks can be pressurised:

a. Complete Pressurisation. Keeping the absolute pressure in the tanks greaterthan the vapour pressure at the maximum fuel temperature likely to beencountered eliminates all losses. However, this means that with gasoline typefuels, a pressure of about 8 psi absolute would have to be maintained at altitudeand the tank would be subjected to a pressure differential of 6.5 psi at 50,000

feet. The disadvantage is that this would involve stronger and heavier tanksand a strengthened structure to hold the tanks.

b. Partial Pressurisation. This prevents all liquid loss and reduces the evaporativeloss. It involves strengthening the tanks and structure and the fitting of reliefvalves.



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9.3.11 FUEL ADDITIVES

Additives are added to fuel to improve its characteristics.

Lubricity Additive. This is added to the fuel to reduce wear in fuel pumps, FCUsetc. when the fuel does not have sufficient lubricating properties of its own.

Ice Inhibitor. Added to fuel to reduce/prevent ice crystals forming in the fuel andsubsequently blocking fuel filters. This additive may also have biocide properties.

Biocide. This is added to the fuel to prevent microbiological growth at the marginsof free water within the aircraft fuel tanks. It can also be used as a shock treatmentif contamination is suspected or as a preventative measure.

9.3.12 SAFETY PRECAUTIONS

Al l fuel wil l burn!

Wide cut fuel is easier to ignite than kerosene.

Strict No Smoking areas should be established around aircraft when any fuelsystem components are removed or fuel tanks are opened. This is important duringrefuelling and tank venting as fuel vapour present in the vent gasses produce anextremely explosive mixture.

Fuel produces a very high static charge when flowing through pipes and meticulouscare must be taken with bonding or grounding of pipes etc. The charge built up isdependant on flow rate, which is exceptionally high during refuel. Care must betaken when draining fuel from a component, as there is a chance of a staticdischarge occurring. Fuel soaked clothing is a great fire risk as the vapours givenoff are combustible.

Fuel can also cause serious damage to the body. It degreases the skin which cancause dermatitis; the additives can increase the damage. Fuel also attacks sensitiveareas of skin causing fuel burns (chemical burn) which can be extremelyuncomfortable and may require hospitalisation. The chance of fuel burns to the skin

is also increased if clothing becomes soaked, because of the proximity and rubbingaction. Wash hands prior to going to the toilet. Eye protection may be required whenentering systems that may contain fuel or fuel vapours. Avoid touching around youreyes if fuel is on your hands, you will only do it once!

Fuel can be harmful if ingested, therefore hands should be thoroughly washed priorto eating.

Spilt fuel on the floor or aircraft skin is very slippery and can even melt the soles ofsome types of shoe. Spills should be mopped up and disposed of in accordancewith company procedures. Fuel spills should not be washed into domestic drains orsewers Spills to grass areas where as there is a chance of the fuel entering and

polluting the water table below ground must be reported.



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9.4 GAS TURBINE OIL PROPERTIES AND SPECIFICATIONS

IntroductionThere are two basic types of lubrication, they are Hydrodynamic (or film) lubrication,where the surfaces concerned are separated by a substantial quantity of oil, andBoundary lubrication, where the oil film may be only a few molecules thick. Beforedescribing the types of lubrication in depth, it is necessary to explain viscosity.

9.4.1 VISCOSITY

The coefficient of viscosity, also known as dynamic viscosity, is a measure of theinternal resistance of a fluid to relative movement, ie. its thickness, or film strength.Viscosity decreases with increase of temperature, the rate depending on theparticular fluid considered. It is important for a lubricating oil that this rate of change

of viscosity is predictable and is as small as possible. The viscosity index (VI) is anempirical number devised to indicate this change of viscosity with temperature, sothan an oil with a high VI is preferable to one with a low VI.

9.4.2 HYDRO-DYNAMICS OR FLUID FILM LUBRICATION

Fluid film lubrication is the most common form of lubrication. It occurs when therubbing surfaces are copiously supplied with oil and there is a relatively thick layerof oil between the surfaces (may be up to 100,000 oil molecules thick). The oil hasthe effect of keeping the two surfaces apart. Under these conditions the coefficientof friction is very small and may be as low as 0.001.

The lubrication of a simple bearing (such as supports a rotating shaft) is a goodexample of fluid film lubrication (see figure 9.2.). The rotating shaft carries oilaround with it by adhesion and successive layers of oil are carried along by fluidfriction. As the shaft rotates it moves off-centre resulting in a narrow wedge of oilwithin which the pressure increases as the wedge narrows. For efficient lubricationthis wedge, and the resulting increase of pressure, is essential as this keeps thesurfaces apart. If this steady pressure increase breaks down, efficient filmlubrication ceases and boundary lubrication occurs.

Lubrication of a Simple Bearing.

Figure 9.2.



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In film lubrication, viscosity is the important factor because it controls the ability ofthe oil to keep the surfaces apart. A shaft revolving at high speed in a bearing mustbe free to carry oil round with it, with as little drag as possible. The rapid movementof one layer of oil slipping over another, with minimum drag, can only be achievedwith a low viscosity oil. As the rotational speed decreases, the rate of deformationof the oil decreases, therefore the drag decreases and consequently an oil of higherviscosity may be needed if it is to be successfully carried round the bearing.

The running temperature of the bearing is equally as important as the speed ofrotation, as it controls the viscosity of the oil to be used. Bearing temperatures mayvary, hence the need for oils with high VIs.

9.4.3 BOUNDARY LUBRICATION

If a shaft carries an appreciable load and rotates very slowly it will not carry roundsufficient oil to give a continuous film and boundary lubrication will occur in whichthe friction is many times greater than in fluid film lubrication.

Boundary lubrication is said to exist when the oil film is exceedingly thin and mayonly consist of a very few layers of molecules. It occurs due to high bearing loads,inadequate viscosity (possibly due to excessive bearing temperatures), oilstarvation or loss of oil pressure. The friction is independent of the viscosity of theoil, but depends on the load and the “oiliness” of the lubricant. When a lubricatingoil reduces the friction in a bearing to a lower value than that given by anotherlubricant of the same viscosity at the same bearing temperature, it is said to have agreater oiliness. It is thought that the reduction in friction is achieved by the fattyacids in the oil combining chemically with the bearing metal to form a “soap” whichgives a boundary layer between the thin oil film and the bearing material to protectthe metals from welding together.

Boundary lubrication is not a desirable phase of lubrication as rupture of the thin filmmeans wear, a very high surface temperature and possible seizure; thereforelubrication is designed to be hydro-dynamic if possible. However, boundarylubrication often occurs during starting conditions and may occur in piston enginesat the end of reciprocating strokes. There is no precise division between boundaryand fluid film lubrication although each is quite distinct in the way in whichlubrication is achieved. In practice both forms occur at some time giving mixed film

lubrication.

9.5 LUBRICATING OILS

General

Viscosity and VI are the factors which decide the lubricant for a particular purpose.The desirable viscosity for a given purpose is decided by bearing loads andclearances, sliding speeds, oil pump capacity, operating temperatures, etc.Therefore, in a lubricating oil specification, the desired viscosity is specified,together with VI and other safeguards to prevent the use of oil, which woulddeteriorate excessively or corrode the engine. Special engine tests are also carried

out in test engines for each main batch of lubricating oil.



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Extreme Pressure LubricantsExtreme pressure lubricants are designed to work under boundary lubricationconditions. Certain chemicals known as extreme pressure (EP) additives (eg.sulphur, chlorine) give the lubricant the necessary quality. They appear to work inthe same way as fatty acids, in that they combine chemically with the surface of thebearing metals.

Additives

Additives are substances added in small quantities to a lubricating oil to give it moredesirable properties.

Additives to lubricating oils are of the following main types:-h. Extreme Pressure, as discussed. They are not in general use except in certain

helicopter applications.

i. Anti-corrosion, which is used to protect some part of the engine.

j. Detergents, which are used in piston engine oils to keep the engine clean.

k. Viscosity Index improvers, which make the VI as large as possible.

l. Pour Point Depressants, that permit oils to flow at lower temperatures than theywere previously able.

m. Anti-foaming additives, that minimise foaming by increasing the surface tensionof the oil.

n. Anti-oxidants, which may be used to reduce the breakdown of the oil due tooxidation.

9.6 TURBINE OILS

Introduction

For lubrication of a high-speed turbine shaft running in contact bearings, an oil withgood boundary lubrication properties and low viscosity is required. Because of thesmall amount of oil in circulation and the high bearing temperatures, good

resistance to oxidation is essential.The earliest gas turbine engines were developed using straight mineral oils, but theoperational requirements for low temperatures either on the ground or at a highaltitude, led to the development of a range of straight mineral oils with viscosity’s farlower than those of conventional aircraft engine oil of that time. Mineral turbine oilsare very rarely used now.



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9.6.1 FIRST GENERATION SYNTHETIC OILS

With the progressive development of the gas turbine engine to provide a higherthrust and compression ratio, mineral oils were found to lack stability and to sufferfrom excessive volatility and thermal degradation at the higher temperatures towhich they were subjected. At this stage, a revolutionary rather than evolutionaryoil development took place concurrently with engine development; lubricating oilsderived by synthesis from naturally occurring organic products found an applicationin gas turbine engines. The first generation of synthetic oils were based on theesters of sebacic acid, principally dioctyl sebacate. As a class these materialsexhibited outstanding properties which made them very suitable as the basis for gasturbine lubricants.

Unlike straight mineral oils, the synthetic oils relied on additives (and in laterformulations on multi-component additive packages) to raise their performance.This was particularly necessary to improve resistance to oxidation and thermaldegradation (important properties which govern long term engine cleanliness).

9.6.2 SECOND GENERATION SYNTHETIC OILS

The introduction of the by-pass or turbo-fan engine raised further problems; in thisengine the by-pass air acts as an insulating blanket and increases heat rejection tothe lubricant. Therefore the requirement arose for an oil with an even greaterresistance to thermal and oxidative stress. Several synthetic oils which meet thisrequirement have been developed. Known as Type 2 lubricants, they are blendedfrom more complex esters and an additive package consisting of anti-oxidants,load-carrying additives, corrosion inhibitors, metal deactivators and foam inhibitors.

9.6.3 THIRD GENERATION SYNTHETIC OILS

Sustained flight at speeds in excess of Mach 1 aggravates the lubricant problem stillfurther as the kinetic heating of the fuel reduces the effectiveness of fuel-cooled oil

coolers. At Mach 2, oil temperatures may reach 260 - 316C, at which levelstandard ester-based oils degrade rapidly. In some military aircraft, Type 1 andType 2 ester oils are still used under these conditions, but at greatly increased oilchange frequencies. This procedure is expensive to operate as ideally the oilshould remain in the engine for full engine life, with only periodic replenishment.

More complex chemicals have been discovered which are more thermally stablethan esters. However, they have various deficiencies such as poor low temperatureproperties or poor steel-on-steel lubricity. All are more expensive than esters.

High temperature lubricants blended from specially developed ester oils, with newadditives to limit oxidation degradation and corrosiveness and of increased loadcarrying ability, appear to offer the most practical solution for lubricating the jetengines in commercial supersonic transport (SST) aircraft. Many firms have beenactive in developing lubricants of this type and, after many submissions, twolubricants have been adopted for the Olympus 593 engines which power the BAC- Aerospatiale Concorde.



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Intentionally Blank



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10 LUBRICATION SYSTEMS

10.1 INTRODUCTION

There is always friction when two surfaces are in contact and moving, for evenapparently smooth surfaces have small undulations, minute projections anddepressions and actually touch at only a comparatively few points. Motion makesthe small projections catch on each other and, even at low speeds when the surfaceas a whole is cool, intense local heat may develop leading to localised welding andsubsequent damage as the two surfaces are torn apart. At higher speeds and overlonger periods, intense heat may develop and cause expansion and subsequentdeformation of the entire surface; in extreme cases large areas may be melted bythe heat, causing the metal surfaces to weld together.

The gas turbine engine is designed to function over a wider environment and underdifferent operating conditions from its piston engine equivalent and therefore speciallubricants have been developed to cope with the following main problems:

a. High rpm compared with piston engines.

b. Cold starting in winter can mean initial bearing temperatures of -54C which

rapidly increases after starting to 232C. Therefore a good viscosity index andadequate cooling are required.

On the other hand, the following advantages can be claimed for the gas turbine:

a. There are fewer bearings and gear trains.

b. Oil does not lubricate any parts directly heated by combustion and therefore oilconsumption is low.

c. There are no reciprocating loads.

d. Bearings are generally of the rolling contact type and therefore only low oilpressures are needed (40 psi is normal).

Turbo-prop engine lubrication requirements are more severe than those of a turbo- jet engine because of the heavily loaded reduction gears and the need for a high-pressure oil supply to operate the propeller pitch control mechanisms. (Forexample, a twin relief valve in the Dart provides 35 psi for engine lubrication and 70

psi, which is fed to the propeller controller and boosted by a further pump to apressure of 600 psi).

10.2 BEARINGS

The early gas turbines employed pressure lubricated plain bearings but it was soonrealised that friction losses were too high and that the provision of adequatelubrication of these bearings over the wide range of temperatures and loadsencountered was more difficult than for piston engine bearings.

As a result, plain bearings were abandoned in favour of the rolling contact type asthe latter offered the following advantages:

a Lower friction at starting and low rpm.

b Less susceptibility to momentary cessation of oil flow.



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c The cooling problem is eased because less heat is generated at high rpm.

d The rotor can be easily aligned.e The bearings can be made fairly small and compact.

f The bearings are relatively lightly loaded because of the absence of powerimpulses.

g Oil of low viscosity may be used to maintain flow under a wide range ofconditions and no oil dilution or pre-heating is necessary.

The main bearings are those which support the turbine and compressor assemblies.In the simplest case (a single spool engine), these usually consist of a roller bearingat the front of the compressor and another in front of the turbine assembly, with a

ball bearing behind the compressor to take the axial thrust on the main shaft.“Squeeze film” main bearings have been introduced to reduce transfer of rotorvibration to the aircraft. In this type of bearing pressure oil is fed to a small annularspace between the bearing outer track and the housing. Figure 10.1. shows that thebearing will therefore “float” in pressure oil, which will damp out much of thevibration. Squeeze film bearings are fitted to the Spey and all subsequent aeroengines produced by Rolls-Royce (1971) Ltd. They have also been fittedretrospectively to existing engines. In addition to the main bearings, lubrication willalso be required for the wheelcase, tacho-generator, CSDU, alternator, starter andfuel pump drives.



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Squeeze Film Bearing.Figure 10.1.



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Single Spool

Twin Spool Turboprop Engine.

Bearing Location Comparison.Figure 10.2.



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10.3 ENGINE LUBRICATION SYSTEMSThere are basically two types of lubrication system at present in use in gas turbineengines:-

a) Recirculatory. In this system, oil is distributed and returned to the oil tank bypumps. There are two types of recirculatory system:-

(i) Pressure relief valve system.

(ii) Full flow system.

b) Expendable. The expendable or total loss system is used on some smallturbo-jet engines, eg. RB 162 in which the oil is spilled overboard after

lubricating the engine.

10.3.1 PRESSURE RELIEF VALVE RE-CIRCULATORY SYSTEM

In the pressure relief valve type of recirculatory lubrication system the flow of oil tothe various bearings is controlled by a relief valve which limits the maximumpressure in the feed line. As the oil pump is directly driven by the engine (by the HPspool in the case of a multi-spool engine), the pressure will rise with spool speed. Above a pre-determined speed the feed oil pressure opens the system relief valveallowing excess oil to spill back to the tank, thus ensuring a constant oil pressure atthe higher engine speeds.

A typical relief valve type of recirculatory lubrication is shown in the figure 10.3.The oil system for a typical turbo-prop engine is similar but, as it supplies thepropeller control system, it is more complicated. The oil supply is usually containedin a combined tank and sump formed as part of the external wheelcase. Oil passesvia the suction filter to the pressure pump, which pumps it through the air-cooled oilcooler to the pressure filter. A pressure regulating valve upstream of the filtercontrols the oil pressure. Both oil pressure and temperature indications aretransmitted to the cockpit. The oil flows through pipes and passages to lubricate themain shaft bearings and wheelcases. The main shaft bearings are normallylubricated by oil jets and some of the heavier loaded gears in the wheelcases arealso provided with oil jets, while the remaining gears and bearings receive splash

lubrication.

An additional relief valve is fitted across the pump in the lubrication system of someengines to return oil to pump inlet if the system becomes blocked.



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A Pressure Relief Valve Lubrication System for a Two Shaft Turbojet.Figure 10.3.



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10.3.2 RECIRCULATORY OIL SYSTEM – FULL FLOW TYPE

The full flow lubrication system is an alternative to the pressure relief valve oilsystem and full flow systems are in use as a means of lubricating many modernhigh power gas turbine engines.

The full flow system is similar in many ways to the pressure relief system justdiscussed – i.e. oil is drawn from a tank by a pump and delivered, via a pressurefilter, to various parts of the engine; the oil is then returned by scavenge pumps, viathe oil cooler to the tank; also, air is separated from the oil by a de-aerator andcentrifugal breather.

The major differences from the pressure relief type of recirculatory system are asfollows:-

The flow of oil to the bearings is determined by the speed of the pressure pump,the size of the oil jets and the pressure in each of the bearing housings.

A metered spill of feed oils is fed back to the tank. This spill is calibrated tomatch the pump output to ensure that the oil flow to the bearings, via the oil jets,is the same at all engine speeds.

The relief valve in this system is set to prevent excessive oil pressure in the feedside of the system.

A filter by-pass is not normally fitted. The pressure drop across the filter issensed by a differential pressure switch, any increase in the pressure difference

being indicated to give advance warning of a blocked filter.

10.3.3 ADVANTAGES OF FULL FLOW LUBRICATION

The advantages of full flow lubrication are those of suitable oil flow and oil pressureat all engine speeds. A study of the graph will reveal a difference in oil pressurebetween the pressure relief system and the full flow system and, it will also showthat the pressure difference continues throughout the speed range of the engines,with a crossover point at cruising speed. The relief valve system provides too muchoil pressure at idle rev/min, but because of the relief valve, the oil pressure is belowoptimum at maximum engine speed. In contrast the pressure provided by the oilpump of a full flow system rises progressively with increased engine speed and is

nearer to the optimum value throughout the rev/min range of the engine.

Comparison of Full Flow and Relief Valve Systems. Figure 10.5.



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Full Flow Oil System ( RR Gem).Figure 10.6.



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10.3.4 EXPENDABLE SYSTEM An expendable system is generally used on small engines running for periods ofshort duration. The advantage of this system is that it is simple, cheap and offersan appreciable saving in weight as it requires no oil cooler, scavenge pumps orfilters. Oil can be fed to the bearing either by a pump or tank pressurisation. Afterlubrication the oil can either be vented overboard through dump pipes or leakedfrom the centre bearing to the rear bearing after which it is flung onto the turbineand burnt.

An Expendable Oil System.

Figure 10.7.



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10.4.2 OIL PUMPSThe oil pumps fitted in a recirculatory system are normally gear-type or Gerotor typepumps. The pumps are usually mounted in a pack containing one pressure pumpand several scavenge pumps. They are driven by a common shaft through theengine gear train.

Gear type pumps (Fig.10.10. ) require suitable machining of the gear teeth, or theprovision of a milled slot in the casting (adjacent to the delivery side of each pump),to prevent pressure locking of the gears.

Gerotor type pumps (Fig.10.11.) use an inner and outer rotor, where the inner rotoris driven by the engine, and the outer rotor which has an extra gear tooth rotates

with it. The inner rotor is eccentric to the outer and it is the stepping of the teeth thatpumps the oil. The pump also requires kidney shaped slots as inlet and outlet ports.

The scavenge pumps have a greater capacity than the pressure pump to ensurecomplete scavenging of the bearings in a dry sump system. Furthermore, air tendsto leak into the bearing housings from the air pressurised seals and this aeration ofthe oil means that the scavenge pumps have to pump an increased oil/air volume. As we saw in the previous paragraph the air is subsequently removed by the de-aerator.

Typical Gear Type Oil Pump.Figure 10.9.



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Gear Type Pump.

Figure 10.10.

Gerotor Type Pump.Figure 10.11.



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10.4.3 OIL COOLING All engines transfer heat to the oil by friction, churning and windage within a bearingchamber or gearbox. It is therefore common practice to fit an oil cooler inrecirculatory oil systems. The cooling medium may be fuel or air and, in someinstances, both fuel-cooled and air-cooled coolers are used.

Some engines which utilise both types of cooler may incorporate an electronicmonitoring system which switches in the air-cooled oil cooler (ACOC) only when it isnecessary. This maintains the ideal oil temperature and improves the overallthermal efficiency.

The fuel-cooled oil cooler (FCOC) has a matrix which is divided into sections by

baffle plates. A large number of tubes convey the fuel through the matrix, the oilbeing directed by the baffle plates in a series of passes across the tubes. Heat istransferred from the oil to the fuel, thus lowering the oil temperature.

The fuel-cooled oil cooler incorporates a bypass valve fitted across the oil inlet andoutlet. The valve operates at a pre-set pressure difference across the cooler andthus prevents engine oil starvation in the event of a blockage. A pressuremaintaining valve is usually located in the feed line of the cooler which ensures thatthe oil pressure is always higher than the fuel pressure. In the event of a coolerinternal fault developing, the oil will leak into the fuel system rather than thepotentially dangerous leakage of fuel into the oil system.

Typical Fuel Cooled Oil Cooler.

Figure 10.12.



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The air-cooled oil cooler is similar to the fuel-cooled type both in construction and inoperation – except, of course, that air replaces the fuel as the cooling agent. Onsome engines, the airflow through the matrix is controlled by a flap valve, which isautomatically operated when the temperature of the return oil rises to a pre-determined value. A turbo-propeller engine may be fitted with an oil cooler thatutilises the external airflow as a cooling medium. This type of cooler incurs a largedrag factor and, as kinetic heating of the air occurs at high forward speeds, it isunsuitable for turbo-jet engines.

10.4.4 PRESSURE FILTER

The pressure oil filter housing contains a wire-wound or mesh, Paper or feltelements and incorporates a by-pass valve. The filter housing can be drainedindependently of the main oil system. This is done through a drain valve in thehousing base. When drained, the filter can be removed for examination, servicing,or replacement, as necessary, without disturbing the rest of the system. Typicalpressure filters are illustrated in figure 10.13.

Wire Wound and Paper Type Oil Filters.Figure 10.13.



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Filters are usually fitted with an impending by-pass indicator. This is usually a redpop out indicator which will pop out and stay out it the differential pressure acrossthe filter element exceeds a predetermined value. This value will be less than theby-pass valve value, to allow the filter to be replaced before the filter goes into by-pass. The pop out usually has a thermal lock on it, which prevents the pop outextending when the oil is cold and thick.

Filter Bowl with Pop Out Indicator.

Figure 10.14.



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10.4.5 LAST CHANCE FILTERSome of the gears in the gearboxes and also the main bearing of the engine arelubricated through oil jets. These jets are usually protected by thread-type oil filters.These are often referred to as last chance filters. You may also find small meshfilters doing this job.

10.4.6 SCAVENGE OIL STRAINERS

When the oil has been distributed to all parts of the engine and has done its job, it isreturned to the oil tank by either gravity or pressure from the scavenge pumps.

Each pump returns the oil from a particular part of the engine and is protected by acoarse filter (or strainer) in the return line. This arrangement protects the pumpgears. It also gives an indication of impending component failure if the strainers areexamined for metal particles during periodical inspection.

Thread Type Last Chance FilterFigure 10.15.



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10.4.7 MAGNETIC CHIP DETECTORMagnetic detectors may be fitted into the oil system at various points to collect andhold ferrous debris. They are normally fitted in gearboxes and in the scavengepump return lines to the tank. The collection of ferrous particles on the chipdetector provides a warning of impending (or incipient) failure of a component.Some detectors are designed so that they can be removed for periodicalexamination without having to drain the oil system; others may be checkedexternally by connecting a suitable test circuit to the plug; finally, some areconnected to a cockpit warning system to give an in-flight indication of failure. Thechip detector (see figure10.15.) fits into a self-sealing housing and has a bayonet-type fitting for easy removal.

10.4.8 DE-AERATOR

We have already noted that air from the bearing sealing system mixes with the oiland causes frothing. If the air is allowed to remain in the oil it may cause alubrication failure. To prevent this, a de-aerating device may be installed; thisremoves air from the oil before the oil is re-circulated round the engine by thepressure pump; the air can be vented to atmosphere via the centrifugal breather.

De-aerators are usually tray types fitted in the oil tank or centrifugal type as aseparate item.

Magnetic Chip Detector.Figure 10.16.



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Centrifugal Breather.Figure 10.17.

10.4.9 CENTRIFUGAL BREATHERWhen the oil/air mixture returns to the tank the air is separated by the de-aeratortray and passes through to the gearbox via a vent line. It carries some of the oil withit in the form of a fine mist. The oil/air mist in the gearbox can then pass to thecentrifugal breather (see figure 10.17). As the vanes of the centrifugal breatherrotate, the oil in the mixture is caught in the vanes and thrown back into thegearbox; the air being vented to atmosphere.

10.4.10 PRESSURE RELIEF VALVE

The pressure relief valve shown in thefigure 10.18. controls the oil pressure atthe pre-set value demanded by the system.

The valve is normally integral with thepump assembly and protects the systemfrom excessive pressure. It is usually aspring-loaded plate-type valve, and can onsome engines provide adjustment ofpressure setting.

Simple Pressure Relief ValveFi ure 10.18.



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It is more usual to find a pressure relief valve that varies the pressure with enginespeed or breather pressure. These valves are usually adjustable but usually onlyeffect the max speed oil pressure see Figure 10.19.

This type of valve uses the oil system breather pressure and an adjustable spring tobalance the oil pressure in the main oil feed line to the engine bearings.

Consider Fig. 10.19. With the engine running, the breather pressure plus the springpush the sliding valve to the left and restrict the pump spill back to return. This isbalanced by the pressure from the main feed line trying to move the slide valve tothe right. Should the pressure in the main feed line fall, the breather pressure andspring will move the slide valve further to the left and restrict the oil spill still further.

This will allow more oil to flow to the system, and the oil pressure in the main feedline will increase. The slide valve will then move to the right, and the oil spill to thereturn will be controlled by the main feed line pressure balancing the spring andbreather pressure.

10.4.11 BY-PASS VALVE

This is similar in construction to the normal pressure relief valve just discussed. It isconnected in the system in such a way that, should the oil cooler or the pressurefilter become blocked (so that the oil flow is restricted), the appropriate by-passvalve will operate to re-route the oil. Although the cooling or the filtering has nowbeen by-passed, oil starvation of the oil bearings is prevented. Pop–out indicators

are used to warn of an impending by-pass.

Pressure Relief Valve That Uses Breather Pressure to Vary Pressure.Figure 10.19



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The oil cooler will usually have a thermal by-pass valve which will by-pass thecooler when the oil is cold, thus ensuring that the oil gets up to running temperaturequickly.

10.5 INDICATIONS AND WARNINGS

Indications and warnings vary from aircraft to aircraft, in both the warnings givenand the priority that they are given.

10.5.1 LOW PRESSURE WARNING LAMP

If the oil pressure drops below the safe operating value for the particular system, apressure-sensing switch will initiate a visual warning; the warning usually consists of

a red or amber lamp switching on in the cockpit accompanied by an audio warning.The sensing switch may be a differential pressure switch which senses the pressuredifference between the feed oil pressure and the return oil pressure or a simplepressure switch. When the pressure or difference falls below a pre-determinedlevel, the switch operates to activate the warning circuit. To reduce the cockpit noiseduring taxiing, the audio warning may be inhibited, as engines are often shut downbefore reaching the stand.

Although this system is simple, its warning factor may not be quick enough toprevent serious damage to the engine. This is due to the fact that the warningpressure must be below the normal oil pressure at idle RPM. This means that theengine could be running for some time with a low oil pressure before the warningoccurs. To overcome this problem multiple pressure switches are used andactivated at differing engine RPM’s. For instance, above 85% RPM the low oilpressure warning will come ‘ON’ at 50 psi, below 85% the warning will come on at20psi.

This is a serious warning and the engine must be shut down as soon as possible.

10.5.2 OIL PRESSURE, TEMPERATURE AND QUANTITY INDICATION

See section 14 engine indications for details of these systems.

10.6 OIL SEALSOil seals have been covered in section 8.

10.7 SERVICING

The engine oil level is usually checked after flight or after an engine run. It is notchecked straight after shut-down, as entrained air will give a false reading. It cannotbe checked accurately if left too long as the oil may run out of the tank into thegearbox. So it is normally checked between 20 minutes and 2 hours or as defined inthe aircraft maintenance manual.

The oil system magnetic chip detectors will be checked at the periodicity defined in

the maintenance schedule. Spectrometric Oil Analysis Program (SOAP) samples ofthe oil may be taken when required.



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Filters are replaced when required by the maintenance schedule or if the pop outindicator is out.



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11 ENGINE FUEL CONTROL SYSTEMS

11.1 INTRODUCTION

The thrust of a turbo jet is controlled by varying the amount of fuel that is burnt inthe combustion system and in order to operate the safe temperature limits, theamount of fuel that is burnt must be governed by the amount of air that is availableat the time. The air supply is dependent upon the RPM of the compressor and thedensity of the air at its inlet, so under a constant set of atmospheric conditions, theRPM of the compressor is an indication of the engine thrust. The pilot has control ofthe fuel flow to the combustion system and is able to select any compressor RPM,between ground idling and the maximum permissible which is required for take offconditions, by the operation of a cockpit lever.

In the normal operational environment of an aircraft engine, atmospheric conditionscan vary over a wide range of air temperatures and pressures resulting in changesof air density at the compressor inlet. A reduction in air density will cause areduction in the amount of air delivered to the combustion system at a selectedRPM, with a consequent increase in the combustion chamber temperature. If thefuel flow is not reduced, a rise in compressor RPM will occur accompanied withoverheating of the combustion and turbine assemblies. An increase in air densitywill result in an increase in the amount of air delivered to the combustion system ata selected RPM and unless the fuel flow is increased, a reduction in RPM will occur.

Changes in air density at the compressor inlet are caused by:-

a) Altitude. The density of the air gets progressively less as the altitude isincreased, therefore less fuel will be required in order to maintain theselected RPM.

b) Forward Speed. The faster the aircraft flies then the faster the air is forcedinto the aircraft intake. A well designed aircraft intake will slow down thisairflow, converting its kinetic energy into pressure energy, so that it arrives atthe compressor inlet at an optimum velocity (0.5Mach) with an increase inpressure and hence density. This is known as Ram Effect and plays animportant part in the performance of a turbo-jet. Within certain limits thegreater the ram effect, the greater the air mass flow and more fuel can beburnt at the selected RPM, producing more thrust.

11.2 PURPOSE OF THE ENGINE FUEL SYSTEM

The purpose of the engine fuel system is to deliver to the combustion system, in areadily combustible form, the correct amount of fuel over the whole operating rangeof the engine, under the control of the pilot.



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Block Diagram of a Fuel Control.(JT9D)Figure 11.1.



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11.3 LAYOUT OF TYPICAL SYSTEM COMPONENTSThe figure 11.1. illustrates the layout of components of a representative fuel system.

11.3.1 AIRCRAFT MOUNTED COMPONENTS

a) Fuel Tanks. Stores sufficient fuel for the aircraft’s designed flight duration.

b) Booster Pump. Ensures a constant supply of fuel at low pressure to the inletof the engine driven HP Fuel Pump.

c) Low Pressure Cock. Isolates the engine fuel system from the aircraft fuelsystem in the event of engine fire or for maintenance.

NOTE: These aircraft mounted components will be dealt with in greater detail duringthe Aircraft Systems Phase.

11.3.2 THE ENGINE LP FUEL SYSTEM

LP Fuel Pump.

Form the LP Cock fuel passes to an engine driven LP Fuel Pump which serves twopurposes:

a. To boost pressure of the fuel to prevent cavitation of the HP pump.

b. To provide means of drawing fuel from the fuel tanks in the event offailure of the fuel boost pump in the tank.

These are normally centrifugal type pumps which will boost pressure in the region of5-10 psi.

Fuel/air heat exchanger.

To reduce the possibility of low temperatures forming ice, in the fuel heating isapplied . Fuel heating is achieved by passing the fuel through a form of radiatorwhich uses hot air (or hot oil) to control and maintain fuel temperature abovefreezing.

LP Fuel Filter.

The filter element may be made of felt, paper or in some cases wire wound. Its

purpose is to prevent foreign particles from entering the engine fuel system. Anindication of the filter ‘clogged’ may be provided on the flightdeck. Not withstandingthis a by-pass will be incorporated to ensure that the fuel supply , albeit possiblycontaminated is still available.

11.3.3 THE ENGINE HP FUEL SYSTEM

HP Pump.

Fuel from the LP Fuel filter passes to the HP pump depending on RPM and FCU inthe region of 600-800 psi. This HP fuel is then fed to the fuel control unit (FCU).



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Fuel Control Unit.

The FCU will meter the engines fuel requirements based upon a given set ofconditions at any given time:

a. Throttle position.

b. Ambient pressure (Pamb)

c. Ambient temperature (T12)

d. HP compressor RPM (N2)

e. Compressor discharge pressure (CDP)

Fuel in excess of that required is returned to the inlet side of the HP pump. Metered

fuel is then fed to the flowmeter via a throttles and HP cock.Thrott le and HP cock.

The fuel control operating levers can be a combined throttle and HP cock lever orseparate levers. The position of the throttle lever determines the power required, theHP shutoff cock controls the supply of fuel from the FCU to the burners, whenclosed the engine will be shut down, when open fuel will be available to the burners.

Fuel Flowmeter.

The fuel flowmeter will measure the amount of fuel being fed to the burners andrelay this information to the flightdeck. A gauge calibrated in either pounds or

kilograms will indicate to the operator how much fuel is being consumed an hour. Asecond window within this gauge may also indicate how much fuel the engine hasconsumed by the engine during the flight.

Fuel/oil Heat Exchanger

Similar to the heat exchanger used to heat the fuel, this heat exchanger will use theHP fuel supply to cool the engine oil.

Pressuris ing and Dump Valve.

From the fuel/oil heat exchanger HP metered fuel passes to the pressurising anddump valve. It function is to:

b. Prevent fuel flowing to the burners during the starting phase until suchtime as fuel pressure is sufficient to give good atomisation of the fuel thusensuring good light-up.

c. Allow sufficient pressure to build up within the Fuel Control Unit (FCU)servo/hydraulic control systems ensuring correct metering of fuel supplyis achieved during starting.

d. Enable a rapid dump of fuel remaining in the pipelines to the burners onshutdown.



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Burners.The type of burners employed will vary with design. Two basic types are in commonuse, atomisers and vaporisers, and their common purpose is to supply fuel in areadily combustible form over the whole operating range of the engine.

11.4 FACTORS GOVERNING FUEL REQUIREMENTS

The factors that determine the quantity of fuel that constitutes ‘the correct amount’to be delivered to the combustion system at any one time are:-

a) The RPM selected.

b) The density of the air at the compressor inlet.

c) The rate at which the engine can accept the fuel into the combustion systemunder conditions of engine acceleration.

11.5 REQUIREMENTS OF THE ENGINE FUEL SYSTEM

a) The selection of the RPM must be under the control of the pilot and the systemmust ensure that the maximum permissible RPM is not exceeded.

b) The fuel must be introduced into the combustion system in a readilycombustible form and the system must be able to automatically adjust thefuel flow to match the air available in order to maintain the selected RPMunder all operating conditions.

11.6 ENGINE FUEL SYSTEM COMPONENTS

In order to achieve its purpose, the engine fuel system will incorporate the followingcomponents:-

a) High pressure fuel pump.

b) Fuel flow-controlling devices.

c) Burners.

11.7 FUEL PUMPS

The type of fuel pump used may vary from one engine type to another and theircommon purpose is to supply the correct amount of fuel to the burners at asufficient rate of flow to ensure operation over the whole range of engine operation.The pump is driven by the engine via a suitable gear train.

11.7.1 FUEL PUMP REQUIREMENTS

Because the fuel flow requirements of an engine running at a constant RPM willvary with changing atmospheric conditions, the fuel pump must be capable ofdelivering fuel at flow rates in excess of the maximum engine demand at anyparticular RPM, eg. its output must be variable independently of its speed ofrotation.

The output of the engine driven fuel pump is dependent on engine RPM andcontrolling signals from various fuel flow controlling devices.



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11.7.3 GEAR-TYPE FUEL PUMPThe gear-type fuel pump (see figure11.3.) is driven from the engine and its output isdirectly proportional to its speed. The fuel flow to the spray nozzles is controlled byre-circulating excess fuel delivery back to inlet. A spill valve, sensitive to thepressure drop across the controlling units in the system, opens and closes asnecessary to increase or decrease the spill.

11.8 FUEL FLOW CONTROL

Control of the fuel flow to the burners is by two main methods:-

a) Manual control by the pilot.

b) Automatic adjustment of fuel flow to correct for basic engine requirements.

(i) Changes in intake pressure.

(ii) Excessive fuel to air ratio during engine acceleration.

(iii) Additional controlling devices as determined by specific enginerequirements.

Gear Type Fuel Pump System.Figure 11.3.



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11.8.1 BASIC FLOW CONTROL SYSTEMPrinciple of Fuel Metering

The flow of a fluid through an orifice (jet) depends on the area of the orifice and thesquare root of the pressure drop across it, ie:-

Fuel Flow = Orifice Area x Pressure Drop

Thus it is possible to vary fuel flow by changing orifice area or the pressure dropacross the orifice. In a fuel system the orifice is variable and is in fact the throttlevalve.

11.8.1.1 Application to Flow Control System

In the flow control system the fuel flow required to give a selected RPM is selectedby throttle area under the control of the pilot (manual control). Compensation forair density variation is superimposed on this selection by the altitude sensing controlunit (pressure drop control unit) varying the pressure difference across the throttlevalve.

11.8.1.2 Control Princ iple

The controlling principle of a flow control system is that a constant throttle pressuredrop is maintained irrespective of throttle area (position) for a given height andspeed.

Principle of Fuel Metering Valve.Figure 11.4.



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11.8.1.3 Principle of Flow Control System (See Figure11.6.)

If however, height and speed change, then the altitude sensing unit will vary thepump output and fuel flow (thus throttle pressure drop) by changing the pumpoutput at constant throttle setting.

Constant Pressure Drop.Figure 11.5.

Principle of Barometric Flow Control.Figure 11.6.



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11.9 HYDRO-MECHANICAL CONTROL UNITS

In hydro-mechanically operated flow control units (FCUs), the method of control isto use servo fuel as a hydraulic fluid to vary fuel flow (eg. by varying pump swash-plate angle). The pressure of the servo fuel is varied by controlling the rate of flowout of an orifice at the end of the servo line; the higher the outflow, the lower will beservo pressure and vice versa. There are two types of variable orifice: the half-ballvalve and the kinetic valve.

11.9.1.1 The Half-Bal l Valve.

In this arrangement, a half-ball on the end of a pivot arm is suspended above thefixed outlet orifice (see figure). Up and down movement of the valve varies servofuel outflow and thus servo pressure and pump output.

11.9.1.2 The Kinetic Valve. Figure 11.8.

A line containing pump output fuel is so placed as to discharge on to the face of theservo outflow orifice and the kinetic energy so produced restricts servo fuel bleed. A blade can be moved downwards to interrupt the high-pressure flow; this reducesthe impact onto the servo orifice, thus causing a greater outflow and a reduction inservo pressure (see figure). The kinetic valve is less prone to dirt blockage than thehalf-ball type, although it is more complex.

Half Ball Valve System.Figure 11.7.



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Simple Flow Control.

The Simple Flow Control Unit (see figure 11.9.) comprises a half-ball valve actingon servo fuel bleed, whose position is determined by the action of an evacuatedcapsule (immersed in P1 air) and a piston subjected to the same pressure drop asthe throttle valve. Fuel from the pump passes at pressure P pump through thethrottle, where it experiences a pressure drop to burner pressure P burner. Theresponse to P1 and throttle variations can now be examined.

Throttle Variations.If the pilot opens the throttle, the throttle orifice area increases, throttle pressuredrop reduces and therefore PPUMP falls, PBURNER rises and the piston moves down,allowing the spring to lower the half-ball valve against the capsule force, increasingservo pressure and pump output. The increased fuel flow increases the throttlepressure drop to its original value, returning the half-ball valve to its sensitive

position.P1 Variations.

If the aircraft climbs, P1 will fall, causing the capsule to expand and raise the half-ball valve against the spring force. Servo pressure will fall, swashplate angle willreduce and fuel pump output will reduce. The reduced flow will cause a reducedthrottle pressure drop.

Thus Simple Flow Control keeps the throttle pressure drop constant, regardless ofthrottle position. At very high altitude the system becomes insensitive and it is notused on large turbo-jets. Nevertheless, it is fitted on the Adour and Dart and hasproved to be a reliable and fairly accurate control unit.

Simple Flow Control.Figure 11.9.



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11.9.3 PROPORTIONAL FLOW CONTROL.The Proportional Flow Control Unit (see figure 11.10.) was designed for use onlarge engines with a wide range of fuel flow. The problem of accurate control overthis wide range was overcome by operating the controlling elements on a proportionof the main flow. The proportion varies over the flow range, so that at low flows ahigh proportion is used for control and at high flows, a smaller proportion. Fuelpasses into the controlling (or secondary) line through a fixed secondary orifice andflows out through another orifice to the LP side of the pump. Secondary flow iscontrolled via the proportioning valve and sensing valve, which maintains an equalpressure drop across the throttle valve and secondary orifice. Servo pressure iscontrolled by a half-ball valve operated by P1 and by secondary pressure.

Throttle Variations.If the throttle is opened, its pressure drop is reduced and the proportioning valvecloses until the pressures across the diaphragm are equalised. Thus secondaryflow and pressure are reduced, the piston drops, the half-ball valve closes andpump stroke increases. The increased fuel flow increases secondary pressure untilthe half-ball valve resumes its sensitive position, but the proportioning valveremains more closed than previously, taking a small proportion of the increasedflow.

Proportional Flow Control.Figure 11.10.



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P1 Variations.Variations in P1 will cause the capsule to expand or contract, thus altering the

position of the half-ball valve and altering fuel flow. This tends to cause rapidchanges in secondary pressure with resultant instability; damping is provided by thesensing valve, which adjusts to control the outflow to LP, thus damping secondarypressure fluctuations. The valve is contoured to operate only over a small range ofpressure drops so that during throttle movements it acts as a fixed orifice.

11.9.4 ACCELERATION CONTROL UNITS

The function of the Acceleration Control Unit (ACU) is to provide surge-freeacceleration during rapid throttle openings. There are two main types of hydro-

mechanical ACU in service.

The Flow Type ACU.With the flow type ACU (see figure 11.11.) all the fuel from the pump passesthrough the unit, which compares fuel flow with compressor outlet pressure (P3),which is proportional to engine speed.

The fuel from the pump passes through an orifice containing a contoured plunger;the pressure drop across the orifice is also sensed across a diaphragm.

When the throttle is opened, the pump moves towards maximum stroke and fuelflow increases. The increased flow through the ACU orifice increases the pressure

drop across it and the diaphragm moves to the right, raising the half ball valve andrestricting pump stroke. The engine now speeds up in response to the limited over-fuelling and P3 rises, compressing the capsule. The plunger servo pressure dropsand the plunger falls until arrested by the increased spring force. The orifice sizeincreases, pressure drop reduces and the diaphragm moves to the left, closing thehalf-ball valve and increasing fuel flow. Fuel flow will increase in direct proportion tothe increase in P3.



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Acceleration Control Using Compressor Discharge Pressure.Figure 11.11.



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The Air Switch.In order to keep the acceleration line close to the surge line, it is necessary tocontrol on “Split P3 air” (a mix of P3/P1) initially and then on full P3 at higher enginespeeds. This is achieved by the air switch (or P1/P3 switch) shown in the figure11.12. At low speeds, P3 passes through a plate valve to P1 and the controlcapsule is operated by reduced, or split P3 until P3 becomes large enough to closethe plate valve and control is then on full P3.

Air Switch.Figure 11.12.

Air SwitchFigure 11.12.



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The dashpot Type ACU.

The dashpot ACU uses two co-axially mounted throttle valves, The inner one ismoved by the pilot, the outer (main) throttle valve will move but is controlled by adashpot which slows the valve movement down to limit the acceleration fuel flow.When closing the throttle the pilot pushes both sleeves in together.

LP

HP

THROTTLOUTLET

THROTTLE CONTROL

THROTTLE SERVO

STEADY

CLOSED INITIAL

FINAL

Dashpot ThrottleFigure 11.13.



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11.10 ENGINE PROTECTION DEVICES

Described below are typical protection devices that will override any excessivedemands made on the engine by the pilot or by the control units.

11.10.1 TOP TEMPERATURE LIMITER.

Turbine gas temperature is measured by thermocouples in the jet pipe. Whenmaximum temperature is reached, these pass a signal to an amplifier, which limitspump stroke by reducing pump servo pressure or moves the throttle valve in serieswith the pilot.

11.10.2 POWER LIMITER.

A power limiter is fitted to some engines to prevent over-stressing due to excessive

compressor outlet pressure during high-speed, low altitude running. The limiter(see figure 11.14) takes the form of a half-ball valve which is opened against aspring force when compressor outlet press (P3) reaches its maximum value. Thehalf-ball valve bleeds off air pressure to the ACU control capsule, thus causing the ACU to reduce pump stroke.

Power Limiter.Figure 11.14.



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11.10.3 OVERSPEED GOVERNOR.The engine is protected against over-speeding by a governor, which, in hydro-mechanical systems, is usually fitted on the fuel pump and acts by bleeding offpump servo fuel when the governed speed is reached. On two-spool engines, thepump is driven from the HP shaft and the LP shaft is protected by either amechanical governor or an electro-mechanical device, again acting through thehydro-mechanical control system. There are two types of pump-driven governors:

11.10.3.1 Centrifugal Governor.

The centrifugal type of governor uses the centrifugal pressure of fuel in radialdrillings in the fuel pump rotor to deflect a diaphragm at maximum speed. The

diaphragm operates on a half-ball valve to reduce pump servo pressure and thuspump stroke. The disadvantage of this type is that it needs to be reset if fuelspecific gravity changes. It is seldom used on modern engines.

Centrifugal Governor Figure 11.15.



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11.11 BURNERS

11.11.1 ATOMISER BURNERS

This type of burner presents the fuel in a finely atomised spray by forcing the fuel topass through a small orifice. The size of the orifice is critical because it mustatomise the fuel effectively over a wide range of fuel flows, from idling to take off

RPM.

Some engines have such a wide range of fuel flow requirements that a single orificeis unable to perform the task effectively unless extremely high fuel pressures areused and to combat this a burner with two different sized orifices are used. Duringlow fuel flow requirements, only the small or primary orifice is supplied with fuel andat higher flow rates both primary and secondary orifices are in operation.

HP Hydro Mechanical Governor.Figure 11.17.

FUEL

PUMP

HP SHAFT

GOVERNOR

ROTATING

SPILL VALVE

LP FUEL IN HP FUEL OUT

SERVO FUEL

LP FUEL

LP FUEL

GOVERNOR FUEL

SERVO FUEL

HP FUEL

HP Hydro-Mechanical Governor.Figure 11.17.



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Both types of atomiser burners incorporate an air shroud, which directs some of theprimary air into the burner to assist atomisation and to cool the burner head to

prevent the formation of carbon.The usual method of atomising the fuel is to pass it through a swirl chamber wheretangentially disposed holes or slots impart swirl to the fuel by converging itspressure energy to kinetic energy. In this state, the fuel passes through thedischarge orifice where the swirl motion is removed as the fuel atomises to form acone-shaped spray. The shape of the spray is an important indication of the degreeof atomisation; thus, the rate of swirl and therefore the pressure of the fuel at theburner are important factors in good atomisation.

Simplex Burner Nozzle Detail.Figure 11.18.



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The simplex burner

The Simplex burner shown in the figure 11.19. was first used on early jet engines. It

consists of a chamber, which induces a swirl into the fuel and a fixed area atomisingorifice. This burner gave good atomisation at the higher fuel flows, that is at thehigher burner pressures, but was very unsatisfactory at the low pressures requiredat low engine speeds and especially at high altitudes. The reason for this is that theSimplex burner was by the nature of its design a “square law” burner, that is theflow through the burner is proportional to the square of the pressure drop across it.This meant that if the minimum pressure for effective atomisation was 30 lbf/in2, thepressure needed to give maximum flow would be about 3,000 lb/in2.

A Simplex Burner.Figure 11.19.



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The Duplex burner.The Duplex burner or Duple burnerrequire a primary and a main fuelmanifold and have two independentorifices, one much smaller than theother. The smaller orifice handlesthe lower flows and the larger orificedeals with the higher flows as theburner pressure increases. Apressurising valve may be employedwith this type of burner to apportionthe fuel to the manifolds (see figure11.20.). As the fuel flow andpressure increase, the pressurisingvalve moves to progressively admitfuel to the main manifold and themain orifices. This gives combinedflow down both manifolds. In thisway, the Duplex and the Dupleburner are able to give effectiveatomisation over a wider flow range

than the Simplex burner for the samemaximum burner pressure. Also,efficient atomisation is obtained atthe low flows that may be required athigh altitude. In the combinedacceleration and speed controlsystem the fuel flow to the burners isapportioned in the FFR. A Duple or Duplex Nozzle.

Figure 11.20.

A Duple or Duplex Burner.Figure 11.20.



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11.11.1.1 The Spray nozzle.The spray nozzle (see figure11.21.) carried a proportion of the primary combustionair with the injected fuel. By aerating the spray, the local fuel-rich concentrationsproduced by other types of burner are avoided, thus giving a reduction in bothcarbon formation and exhaust smoke. An additional advantage of the spray nozzleis that the low pressures required for atomisation of the fuel permits the use of thecomparatively lighter gear-type pump.

A Spray Nozzle.Figure 11.21.



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11.11.2 VAPORISING BURNERSThis type of burner presents the fuel in the combustion system in the form of a richfuel vapour or gas. This is achieved by delivering the metered flow of fuel to “J”shaped vaporising tubes, which protrude into the combustion chamber. The fuelpasses down the vaporising tubes in a coarse spray and mixes with the primary airthat enters concentrically to the fuel supply pipe. The fuel and air is mixedthoroughly by pins that protrude into the primary airflow and the heat of the flamesurrounding the tube causes the mixture to vaporise before it emerges in thecombustion chamber.

The introduction of the primary air into the vaporising tubes aids the process of

vaporisation and also helps to cool the tubes to prevent the formation of carbon.With this type of burner, the flame points towards the incoming airflow and thishelps to stabilise the flame in the vaporising tubes, preventing it being blown awayby the secondary air, thus allowing a relatively short combustion system.

The advantages of this type are:-

a) Pre-vaporising gives complete combustion within a short length of flame tube.

b) A complete ring of flame around the annular chamber.

c) Even pressure and temperature around the chamber.

A Vaporising Combustion Chamber.Figure 11.22.



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More power and efficiency result from “rich” mixtures, but these are limited bymaximum turbine temperatures. Therefore fuel supplies must be limited so that anoverall air/fuel ratio of about 60:1 at maximum rpm is achieved. At other rpm theratio will change due to changing efficiencies of turbine and compressor. The“correct” mixture strength is 15:1 hence only about a quarter of the air passingthrough the engine is used for combustion. (15% - 25% is the typical range).

In the flame area the ratio is about 13:1 and around the flame centre a weaker ratioof 18:1 is used to ensure complete combustion with no carbon formation.

The flame rate at an atomising burner is 2-10 ft/sec and at a vaporiser, 60 ft/sec.Both figures are low compared with the air velocity through the combustion zone,hence the requirement for a low velocity zone at the burner to (a) aid ignition and (b)maintain the flame at the burner.

Theoretically, combustion in a gas turbine is at “constant pressure”, ie. the pressurealong the combustion chamber does not change due to combustion but could alterdue to changes in rpm and air intake pressure.

In practice the combustion chamber shape affects the pressure and they aredesigned to minimise this and a drop of 4% along its length is usual.

Flame temperature is high; a constant 2,000C at the centre. Flame size, however,can change and the bigger the flame becomes the higher goes Turbine EntryTemperature and Jet Pipe Temperature (TET and JPT).

“Over-fuelling” gives a larger flame and “Under-fuelling” a smaller; the significanceof these will be seen in a later note.



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11.12 ELECTRONIC ENGINE CONTROL SYSTEMS

Advances in gas turbine technology have demanded more precise control of engineparameters than can be provided by hydromechanical fuel controls alone. Thesedemands are met by electronic engine controls, or EEC, of which there are twotypes: supervisory and full-authority.

11.12.1 SUPERVISORY ELECTRONIC ENGINE CONTROL

The first type of EEC is a supervisory control that works with a provenhydromechanical fuel control.

The major components in the supervisory control system include the electroniccontrol itself, the hydromechanical fuel control on the engine, and the bleed air andvariable stator vane control. The hydromechanical element controls the basic

operation of the engine including starting, acceleration, deceleration, andshutdown. High-pressure rotor speed (N2), compressor stator vane angles, andengine bleed system are also controlled hydromechanically. The EEC, acting in asupervisory capacity, modulates the engine fuel flow to maintain the designatedthrust. The pilot simply moves the throttle lever to a desired thrust setting positionsuch as full takeoff thrust, or maximum climb. The EEC adjusts the fuel flow asrequired to maintain the thrust compensating for changes in flight andenvironmental conditions. The EEC control also limits engine operating speed andtemperature, ensuring safe operation throughout the flight envelope.

If a problem develops, control automatically reverts to the hydromechanical system,

with no discontinuity in thrust. A warning signal is displayed in the cockpit, but noimmediate action is required by the pilot. The pilot can also revert to thehydromechanical control at any time.

Electronic Engine Control

A typical example of an EEC system is that used in many of the Pratt and Whitney100 series engines currently in service. A brief explanation of how the systemworks, both in automatic and manual modes follows.



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Pratt & Whitney 100 Series Fuel Control System Schematic.Figure 20.24.



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Automat ic Operation (EEC mode)The EEC receives signals from various sources:

a. Power Management Switch, enabling take off thrust, maximum continuousthrust, climb thrust or cruise thrust settings to be selected.

b. Engine inlet pressure and temperature.

c. Ambient pressure.

d. Air data computer inputs. (a computer that senses pitot pressure, static pressureand total air temperature)

e. Engine RPMs – N1 and N2.f. Power lever position. (via a potentiometer)

g. Failure signals.

Based on these input signals the EEC will output command signals to adjust andcontrol:

a. The Hydromechanical Fuel Control Unit via a stepper motor which adjusts thethrottle metering valve.

b. Ignition circuits.

c. Bleed valves

d. Torque gauge

11.12.2 FUEL CONTROL

11.12.3 GENERAL

The fuel control is provided by the hydro-mechanical unit (HMU) The HMU issupplied by the HP fuel pump and provides the required fuel quantity to thenozzles.

In normal operation the fuel control is managed by the Electronic EngineControl (EEC). This enables accelerations and decelerations without engine surge

or flame out whatever the displacement sequence of the power lever. The HMU isalso mechanically connected to the power lever thus ensuring fuel control in caseof failure of the EEC.

Hydro-mechanical Unit (HMU)

The HMU comprises:

A stepper motor controlled by the EEC.

A lever which controls fuel shutoff.

A lever which controls the fuel flow.



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PW100 Series Fuel System Auto/NormalFigure 20.25.



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Operation

The fuel flow supplied to the nozzles is mainly obtained through two valves:a bypass valve

a metering valve.

The fuel enters the HMU from pump outlet with a constant flow. This flow is split bythe bypass valve into two flows, one for the nozzles (via the metering valve) and onebypass return flow to the pump. The position of the bypass valve is a function of theloss of fuel pressure caused by the metering valve. The metering valve ispneumatically actuated. In the pneumatic servo block, the reference pressure is theHP compressor outlet pressure, P3. A controlled reduction of the P3 pressure resultsin a variable Py pressure which when opposed to a bellows device, moves the pistonof the metering valve.

The pneumatic servo block is managed:

in normal operation by the EEC

in manual operation, by the power input lever.

Normal Operation (EEC Mode)

According to the input data (pressures, temperatures, speeds) and to thecommanded power (power lever), the EEC controls a stepper motor located in theHMU.

The stepper motor regulates Py pressure thus modulating the fuel flow asrequested. A governor acts on the Py pressure, thus setting an NH speed limitfunction of the compression of a spring by a cam (EEC cam) connected to thepower lever.

Manual Operation (Manual Mode)

Py pressure is not regulated by the stepper motor but by the simultaneousactions of the NH speed governor and the spring, compressed by a second cam(manual cam) connected to the power lever.

Transfer from the EEC Mode to the Manual Mode.

In normal operation the EEC manages the fuel regulation. The manualoperation is automatically connected when the operation in the EEC mode isswitched off. A solenoid in the HMU selects the manual cam instead of the EECcam and cancels the regulation control through the stepper motor.

Operation of the HMU in the fail mode

In case of failure of the EEC, the position of the stepper motor is "frozen".Whatever the increase of power through the power lever, the last NH speedremains unchanged (the load applied by the spring on the NH speed governorincreases).For any power reduction through the power lever, the NH speeddecreases according to the curve of the EEC cam (decreasing spring load).



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PW 100 Series Fuel System in Manual Mode.Figure 11.26.



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Fuel Distribut ionDuring operation, fuel flows from the aircraft fuel tank to the fuel-pump boost-stageinlet. The pressurised fuel from the boost stage of the engine-driven fuel pump thenleaves the pump and is delivered to the fuel/oil cooler, whose purpose is to keep thefuel sufficiently warm to prevent ice from forming in the fuel, and at the same time,keep the maximum temperature of the oil within the correct limits. This engine isalso equipped with an air/oil heat exchanger, which uses fan air and 2.5 bleed air toprevent the fuel from getting too hot.

From the fuel/oil cooler, the fuel is returned to the fuel pump, where it is filtered andsent to the main pump stage to be further pressurised before it is sent to the fuel-

metering unit, which actually does the metering on the basis of information itreceives from the FADEC. The fuel-metering unit sends fuel to the fuel-flowtransmitter, and then to the fuel distribution valve. (Servo fuel, used as an actuationpressure to some interface components, also comes from the fuel-metering unit.)Bypass fuel not sent to the fuel distribution valve or servo supply is returned topump interstage flow. From the fuel distribution valve, the metered fuel flowsthrough the fuel manifolds to the fuel injectors.

The FADEC is the primary interface between the engine and the aircraft. TheFADEC contains two channels that are called "A" channel and "B" channel. Eachtime the engine starts, alternate channels will automatically be selected. The

channels are linked together by an internal mating connector for crosstalk datatransmission. Much more is accomplished by this control than simply sending asignal to the fuel-metering unit to establish a fuel flow to the nozzles.



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Interface with Aircraft

The FADEC receives several refereed (a validated reference used to confirmcorrect input) inputs and delivers several outputs. Inputs to the FADEC come fromthe following:

1. The power levers. Two analogue signals come from each power-lever resolver.(The resolver is an electromechanical device to measure angular movement.)

2. The air-data computers (ADC) in the form of

a. Total pressure

b. Pressure altitude

c. Total air temperature

3. The flight-control computer (FCC) for adjusting the engine pressure ratio (EPR)for all engines as a part of the engine thrust trim system (ETTS). The ETTSlogic starts when the engine pressure ratio (EPR) on any two engines is above1.2.

4. Seven discrete (electrical signals) inputs:

a. Pt2/Tt2 probe heat

b. Fire

c. Alternate mode select

c. External reset (fuel-control switch)

FADEC Interface with the Aircraft.Figure 11.28.



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d. Bump rate selector

e. Maintenance (data retrieval)f. Engine location identification

5. Two sources of 28 VDC power (DC bus and ground test power)

Out puts from the FADEC are as follows:

Engine pressure ratio (EPR)

Low-speed spool (NI). There is a backup N1 speed output from channel "B."

Exhaust gas temperature (EGT)

High-speed spool (N2)

Flap/slat position and weight-on-wheels status is also sent to the FADEC. Theflight-control computer (FCC) acts as a backup for the air-data computer (ADC).

FADEC Interface wi th Engine

All data input to the FADEC is validated through a series

of comparisons and checks .For example, compressor rotor speeds are comparedto each other and checked to ensure the proper range (0 -120 percent).

Inputs to the FADEC from the engine are as follows:

N2 rpm, Power comes from the FADEC alternator and is used for limiting,

scheduling systems, and setting engine speeds.

N1 rpm, which comes from the FADEC speed transducer (a transducer is adevice used to transform a pneumatic signal to an electrical one) and is used forlimiting and scheduling systems. It is also used as an alternate mode.

Compressor-exit temperature (Tt3), which comes from the diffuser case, is usedto calculate starting fuel flow. • Exhaust-gas temperature (Tt4.95), which comesfrom the exhaust case, is used for indication.

Fuel temperature (Tfuel), which comes from the fuel pump, is used to schedulethe fuel heat-management system.

Oil temperature (Toil), which comes from the main gearbox, is used to schedulethe fuel heat-management system and to schedule the integrated drive generator(IDG) oil-cooling system.

Inlet total temperature (Tt2), which comes from the inlet cowl on the wingengines and the bellmouth on the tail engine. It is used to calculate fuel flow androtor speed.

Inlet total pressure (Pt2), which comes from the same sources as Tt2, is used tocalculate EPR.

Exhaust gas pressure (Pt4.95), which comes from the exhaust case, is also used

to calculate EPR.



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The engine electronic control (EEC) programming plug is used to determine theengine thrust rating and EPR correction.

Burner pressure (Pb), which comes from the diffuser case, is used for limiting andsurge detection. • Ambient pressure (Pamb), which comes from the inlet cowl, isused to validate altitude and Pt2.

Based on information received from its various sources the FADEC will:

1. Monitor, control and protect:

Anti surge bleed valves/variable stator vanes

Cooling airflows

Engine oil cooling and IDG oil cooling

Nacelle cooling

Fuel heating

Starting

Idle speed Acceleration/Deceleration

FADEC Interface With Engine.Figure 11.29.



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Stabilised engine operation

Thrust control including overboost

Critical speeds and pressures

2. Improve reliability of the engine by:

A two channel system of control

An automatic fault detection and logic system

An automatic fault and compensation system

3. Make maintenance easier by:

Engine monitoring

Self test

Fault isolation

Control Modes

The FADEC has two modes for setting the power of the engine. The EPR mode isthe rated or normal mode, while the N1 mode is the alternate or fault mode.

Normal Mode. When a thrust-level request is made through the thrust lever, thethrottle-resolver angle (TRA), input causes an EPR command. The FADEC will thenadjust fuel flow so that EPR actual equals EPR command.

The normal or rated power levels are

Maximum power available (takeoff or maximum continuous)

Maximum climb

At approximately 78 degrees TRA maximum power available is calculated by theFADEC. If the altitude is less than approximately 14,100 ft, the FADEC calculates atakeoff power rating. But if the altitude is greater than 14,100 ft, the FADECcalculates a rating for maximum continuous power. At approximately 68 degreesTRA, the FADEC calculates the maximum climb-power rating. To get all other powerlevels, except idle, it is necessary to set the thrust lever.

Al ternate or N1 Mode.

If the FADEC cannot control in the EPR, or normal mode, it will go to the N1 modeand a fault is enunciated . In the N1 mode, the FADEC schedules fuel flow as afunction of the thrust-lever position, and the TRA input will cause the FADEC tocalculate an N1 command biased by Mach number, altitude, and Tt2. In reversethrust, the FADEC goes to the N1 mode, and N1 is biased by Tt2.

Control in the N1 mode is similar to that of a hydromechanical fuel-control system.Moving the thrust lever fully forward will cause an overboost of the engine.

N1 mode may be manually selected, but the logic that keeps the thrust at the samelevel as it would be in the EPR mode is removed.



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Faults

The FADEC has dual electronic channels, each with its own processor, powersupply, program memory, selected input sensors, and output actuators. Power toeach electronic control channel is provided by a dedicated, engine gearbox-drivenalternator. This redundancy provides high operational reliability. No single electronicmalfunction will cause an engine operational problem. Each control channelincorporates fault identification, isolation, and accommodation logic.

Parameters Sensed and Controls Actuated by an Electronic Engine Control.

Figure 11.30.



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While electronic controls are highly reliable, malfunctions can occur. A hierarchy offault-tolerance logic will take care of any single or multiple faults. The logic alsoidentifies the controlling channel, and if computational capability is lost in the primarychannel, the FADEC automatically switches to the secondary channel. If a sensor islost in the primary channel, the secondary channel will supply the information. If datafrom the secondary channel is lost, the FADEC will produce usable synthesisedinformation from the parameters that are available. If there is not enough dataavailable for synthesising, the control modes switch. For example, if EPR is lost, theengine will be run on its N1 ratings.

In the unlikely event both channels of electronic control are lost, the torque motorsare spring-loaded to their fail-safe positions. The fuel flow will go to minimum flow,the stator vanes will move to fully open, the air-oil cooler will open wide, and the

ACC will shut off.

The FADEC includes extensive self-test routines which are continuously actuated.BITE, or built-in test equipment, can detect and isolate faults within the EEC and itsinput and output devices. The fault words of the control are decoded into Englishmessages by a maintenance monitor, and they identify the faulty line-replaceableunit (LRU). In-flight fault data is recorded so it can be recalled during shop repair.The FADEC is able to isolate problems and indicate whether the fault is within itselfor in a sensor or actuator. In the shop, computer-aided troubleshooting can identifya fault at the circuit-board level.

EEC Programming Plug

The EEC programming plug located on the FADEC "A" channel housing, selects theapplicable schedules within the FADEC for the following:

Engine thrust rating

EPR modification data

Engine performance package

Variable-stator-vane schedule

2.9 bleed-valve thermocouple selection

The EEC programming plug data is input to the FADEC "A" channel, while the "B"channel EEC programming-plug input is crosswired and crosstalked from the "A"channel. During test-cell operation, the EPR/thrust relationship is compared, and theengine gets a correct EEC programming plug. If the FADEC must be replaced, theEEC programming plug must remain with the engine.

If the engine is started without the EEC programming plug installed, the FADECgoes to the N1 mode. But nothing will happen with the FADEC operation if the EECprogramming plug disconnects in flight.



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EEC Programming Plug.Figure 11.31.



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Pneumatic and Electrical Connectors As shown in Figs:11.32. there are several pneumatic and electrical connectors tothe FADEC. The four pneumatic inputs are as follows:

1. Pt4.95 This input comes from two combination Pt4.95/Tt4.95 probes, located onthe turbine exhaust case, and goes to FADEC port "P5." For all pressure inputs atransducer in the FADEC changes the pressure signal into an electric signal andsends this signal to both channels.

2. Pt2 This input comes from the Pt2/Tt2 probe located in the inlet duct.

3. Pb This input comes from a static pressure port in the diffuser case to measure

burner pressure.4. Pam-This input comes from two screened static pressure ports located on the inlet

cowl outer surface.

FADEC Electrical and Pneumatic Connections.

Figure 11.32.



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Al ternator .The alternator provides the FADEC with power and an N2 speed signal. It alsosends N2 information to the flight deck.

FADEC Alternator Figure 11.33.



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Speed Transducer. The speed transducer supplies the FADEC "A" and "B" channelswith the N1 signal by sensing the frequency at which the 60 teeth on the low-pressure compressor/low-pressure turbine (LPC/LPT) coupling pass by them.

Temperature Probes.

A dual-element, alumel-chromel thermocouple, located on the top right side of thefuel pump, provides the FADEC with information relating to fuel heating and engineoil cooling. Oil Temperature Probes. Two other similar devices inform the FADECabout scavenge oil temperature and No. 3 bearing-oil temperature, and provideinput for engine oil cooling-system control, oil-temperature warning indication, andIDG oil-cooling override.

Tt3 Temperature Probe.

This dual-element probe is located on the diffuser case and provides the FADECwith information for heat-soaked engine start logic.

FADEC Speed Transducer Figure 11.34.



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Tt14.95 Temperature Probes.

Four thermocouples measure EGT and send their signal to the thermocouple junction box and then to the FADEC. The temperature sense is used only for inputto the indication system. There is no EGT limiting function in the FADEC.

Exhaust Gas Pressure Probes.

The two probes measure Pt14.95 pressure, are manifolded together, and send theiraveraged pressure to the FADEC.

FADEC Fuel and Oil Temperature Thermocouples.Figure 20.35.



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FADEC Exhaust Gas Temperature and Pressure Probes.Figure 11.37.

FADEC T6 Probe and Exhaust Gas Temperature Junction BoxFigure 11.36.



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Automat ic Turbine Rotor Clearance Control SystemThe automatic turbine rotor clearance control system also known as the turbine casecooling system, controls and distributes fan air to cool and shrink the HPT and LPTcases. This process increases efficiency by reducing turbine tip clearance fortakeoff, climb, and cruise operation. The FADEC commands the system operation toa schedule determined by altitude and N2.

Turbine Vane and Blade Cooling System

The turbine vane and blade cooling system (TVBCS) optimises engine performanceduring cruise by controlling 12th-stage cooling airflow to the HPT and LPT areas.This system is also controlled by the FADEC as a function of altitude and N2.

Additionally, the FADEC receives a feedback signal from the TVBCS right valve.

Turbine Case Cooling System.Figure 11.39.



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FADEC Controlled Active Tip Clearance System

Figure 11.40.



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Turbine Vane and Blade Cooling System.Figure 11.41.



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A Pressure Control System for a Turbo –Prop Engine (Dart)Figure 11.42



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Figure 11.43. A Pressure Control System for a Turbo-Jet Engine (Adour).



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A Proportional Flow Control System (Avon).Figure 11.44.



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Combined Acceleration andSpeed Control.(Spey & Tay).

Figure 11.45.



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Combined Speed and Acceleration Control with Air Bleed Control. ( ALF502.)

Figure 11.46.



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texto turbinas a gas 2013 001

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