5. ENGINE CONTROLS 6 - · PDF fileThe flight compartment throttle lever movement is transmitted to the FCU which in turn controls engine speed in one of the following two modes:

OPERATING MANUAL PSP 601A-6

SECTION 17

POWER PLANT

TABLE OF CONTENTS

Page

1 . GENERAL 1

2. ENGINE FUEL SYSTEM 1

A. Firewall Fuel Shutoff Valve 2 B. Fuel Pressure Sensor 2 C. Engine Fuel Pump 2 D. Fuel Heater ( A i r c r a f t 5001 to 5134) 2 E. Heat Exchanger (A i rc ra f t 5135 and subsequent) 2 F. Fuel Temperature Sensor 2 G. Fuel Filter 2 H. Fuel Control Unit (FCU) 3 I . Fuel Flow Transmitter 3 J. Oil Cooler 3 K. Fuel-Row Distributor and Injectors (Aircraft 5001 to 5134) 3 L. Fuel Manifold (Aircraft 5135 and subsequent) 3 M. Fuel Injectors (Aircraft 5135 and subsequent) 3 N. Ecological Drain System (Aircraft 5001 to 5134) 3

3. AUTOMATIC PERFORMANCE RESERVE (APR) SYSTEM 4

4. ENGINE OIL SYSTEM 4

A. Oil Replenishment System 4 B. Oil Storage Tank 5 C. Oil Temperature Sensor 5 D. Oil Circulation 5 E. Oil Filter 5 F. Oil Cooler 5 G. Oil Pressure Sensor and Low Oil Pressure Switch 5

5 . ENGINE CONTROLS 6

6. THRUST REVERSER 6

6 7 7 7 7 8 8

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A. B.

Operation Safety Features (1) Throttle Retarder System (2) Throttle Lockout System (3) Auto Slow System (4) Emergency Stow System (5) Safety relay


Page

7. ENGINE INSTRUMENTS 8

A. Signal Data Converter (SDC) 8

B. Engine Instruments 8

8. EN6INE BLEED AIR 8

A. Tenth Stage Bleed Air 9 B. Fourteenth Stage Bleed Air 9

C. Bleed Air Leak Detection and Warning System 9

9. ENGINE STARTING AND IGNITION SYSTEMS 10

A. Ground Starting 10 B. In-Right Starts 10 C. Continuous Ignition 11

10. ENGINE VIBRATION MONITORING SYSTEM 11

LIST OF ILLUSTRATIONS

Figure Title Page Number

1 Power Plant - Schematic 12

2 Engine Fuel System - Schematic (2 Sheets) 13

3 Fuel Control Panel - Engine Fuel System Monitoring 14A

4 APR and Engine Speed Control Panel 15

5 Engine Oil System - Schematic 16

6 Oil Temperature and Pressure Indicators 17

7 Throttle Quadrant and Thrust Reverser Controls and Indicators 18

8 Thrust Reverser Stowed and Deployed Positions 19

9 Engine Instruments and Control Panel 20

10 Tenth Stage Engine Bleed Air - Schematic 21

11 Bleed Air Control Panel 22

12 Fourteenth Stage Engine Bleed Air - Schematic 23

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cfianencjer OPERATING MANUAL

PSP 601A-6

Figure

Number Title Page

13 Bleed Air Leak Warning and Testing 24

14 Engine Start and Ignition Controls 25

15 Engine Vibration Monitor Panel 26

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SECTION 17

POWER PLANT

. GENERAL (Figure 1)

The aircraft is powered by two General Electric CF34 turbofan engines. The engine is a dual-rotor, front-fan configuration with a bypass ratio of: 6.2:1 (aircraft 5001 to 5134) 6.26:1 (aircraft 5135 and subsequent). The low pressure (or NJ rotor consists of a single-stage fan driven by a four-stage low-pressure turbine. A high-pressure (or N2) rotor consists of a fourteen-stage axial-flow compressor driven by a two-stage turbine. For a two-engine operation under standard sea-level conditions, the engine is rated at a take-off thrust of 8,729 pounds. An automatic performance reserve (APR) system increases the standard take-off thrust rating to 9,220 pounds if an engine failure occurs.

The engine airflow passes through the fan assembly and is divided into two airflow systems. The main airflow, bypass air, is routed around the core cowls and exhausts through the thrust reverser assembly, over the tailpipe fairing. The remaining airflow passes through the engine core consisting of the compressor, a combustion chamber, the high-pressure turbine and the low-pressure turbine. The hot gas is then exhausted through an exhaust nozzle.

The compressor has a variable geometry system that varies the position of the compressor inlet guide vanes and the first five stages of the stator vanes. The system operates throughout the operating range of the engine to improve compressor efficiency and prevent stalling and surging.

An accessory gearbox, driven by the N2 rotor, drives the engine lubrication pumps and fuel pump as well as an aircraft hydraulic pump and ac generator. The engine starter drives the N2 rotor through this accessory gear box. Bleed air is taken from the 10th and 14th stages of the compressor for air conditioning/pressurization, engine crossbleed starting, anti-icing and thrust reverser operation.

. ENGINE FUEL SYSTEM (Figures 2 and 3)

Each engine has a self-contained fuel system for the controlled distribution of fuel to the combustion chamber. Secondary functions of the system are control of the compressor variable geometry system, cooling of engine oil, and motive fuel supply to the aircraft fuel system ejector pumps (refer to Section 12).

Sensors are installed at suitable locations in the system to provide the required inputs to the flight compartment controls and indicators.

On aircraft 5001 to 5134, the principal components that make up the engine fuel system are described in a fuel flow sequence, starting at the firewall fuel shutoff valve through to the combustion chamber and the ecological drain system.

On aircraft 5135 and subsequent, the principal components that make up the engine fuel system are described in a fuel flow sequence* starting at the firewall fuel shutoff valve through to the combustion chamber.

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esnEmSntyttr OPERATING MANUAL

PSP 601A-6

A. Firewall Fuel Shutoff Valve

This valve isolates the engine fuel system from the engine feed line (refer to Sections 9 and 12).

B. Fuel Pressure Sensor

This sensor, connected to a warning light in the flight compartment, allows monitoring of fuel pressure from associated ejector and electric fuel pumps in the aircraft fuel system,

C. Engine Fuel Pump

The engine-driven fuel pump contains a low-pressure section and a high-pressure section. The low-pressure section supplies fuel through the fuel heater to the fuel f i lter, and back to the high-pressure section of the pump. The high-pressure section is divided into two elements, designated the primary element and the secondary element. The primary element pumps fuel from the fuel filter to the fuel control unit and the secondary

I element supplies high-pressure fuel to the aircraft tanks for motive I flow. I D. Fuel Heater (Aircraft 5001 to 5134)

The fuel heater is an air-to-liquid heat exchanger which uses hot compressor bleed air to heat the fuel. The fuel temperature is maintained above 5*C, by a thermal sensor and an air modulating valve, to prevent icing in the fuel filter. A fuel bypass valve allows fuel to bypass the fuel heater should i t become clogged.

| E. Heat Exchanger (Aircraft 5135 and subsequent)

I Fuel is heated by the liquid to liquid (oil to fuel) heat exchanger.

I F. Fuel Temperature Sensor

This sensor, connected to an indicator in the flight compartment, allows monitoring of the fuel temperature and the fuel heater operation.

| 6. Fuel Filter

The fuel filter, located downstream of the fuel heater, contains a disposable filter element. A bypass valve allows the fuel to bypass the filter element should it become clogged. A differential pressure switch connected to a warning light in the flight compartment warns of an impending bypass condition. Should a bypass occur, a red button on the housing rises to indicate the condition.

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J H. Fuel Control Unit (FCU) The FCU is a hydro-mechanical/electrical device consisting of two sections: a fuel metering section and a computer section. The flight compartment throttle lever movement i s transmitted to the FCU which in turn controls engine speed in one of the following two modes: - At relatively low power settings, the FCU hydro-mechanically meters the

fuel to the injectors to control engine N2 speed. In this mode, matched movement of the throttle levers produces matched N2 speeds, but Ni speeds and thrust may be mismatched between the engines.

- At take-off, climb and cruise power settings, the engine is Ni speed controlled. In this mode, the FCU electrically responds to Ni speed references. (Electrical power for operation in this mode is supplied by an Ni driven alternator, completely independent of the aircraft electrical system). Matched movement of the throttle levers produces matched Ni speeds, hence matched thrust between the engines. This engine speed control can be selected on or off by a switch in the flight compartment. If selected off, the engine speed control reverts to the N2 mode described above for all engine speeds.

The FCU also meters pressurized fuel from the engine-driven fuel pump to the two actuators for the compressor variable geometry system. The actuators move the compressor inlet guide vanes, and the affected stator vanes open as engine speed increases and close as speed decreases.

j I. Fuel Flow Transmitter This transmitter sends fuel flow information to be displayed on the associated fuel flow indicator in the flight compartment.

I J. Oil Cooler The oil cooler heats the fuel before the fuel enters the combustion chamber while cooling the engine o i l . (Refer to paragraph 4.F.)

I K. Fuel-Flow Distributor and Injectors (Aircraft 5001 to 5134) The fuel-flow distributor precisely meters the fuel to the injectors. The injectors inject fuel into the combustion chamber. At shutdown, the distributor drains fuel to an ecological drain system.

J L. Fuel Manifold (Aircraft 5135 and subsequent) Metered fuel leaves the FCU, thru fuel flow transmitter and enters fuel

I manifold. The manifold consists of 2 separate halves, which form one J continuous ring which encircles the combustion chamber frames.

M. Fuel Injectors (Aircraft 5135 and subsequent) Integral with the continuous ring are 18 fuel injector hoses which connect to 18 fuel injectors. Each injector has 2 independent fuel flow passages, a primary and a secondary. The primary introduces fuel into

I the combuster at lower power settings (start-up-idle) the secondary introduces fuel at higher power settings, resulting in 2 cones of fuel.

I N. Ecological Drain System (Aircraft 5001 to 5134) This system prevents the fuel collected by the shut-down fuel drain system from being discharged to the atmosphere. The ecological drain tank collects this fuel which is then routed to an aspirator in the engine fuel feed line to be consumed during the subsequent engine operation.

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3. AUTOMATIC PERFORMANCE RESERVE (APR) SYSTEM (Figures 2 and 4)

The APR system monitors engine thrust levels at high power settings and automatically commands an increase in thrust on both engines if a predetermined thrust loss is detected on one of them. To arm the system, both engine speed control switches must be selected to ON and the APR switch selected to ARM.

With the system armed and the engines operating in the Ni speed control mode, an Ni drop to 5000 rpm (approximately 67.5% Ni) on either engine causes the APR controller to command an N2 speed increase of 167 rpm (approximately 2% Ni) on both engines. The engine st i l l operating at the normal take-off Nx has its Ni increased by approximately 2% while the engine affected by the Ni drop reverts to N2 speed control mode, hence not responding to the Ni speed increase command.

NOTE: The APR system does not affect or override the throttle lever inputs to the FCU. Therefore, it is possible to advance the throttles and obtain power settings higher than the normal (non-APR) take-off thrust. Should this condition be followed by a power loss on one of the engines, the other engine would respond to the APR command and further increase Ni above its previously set higher power setting, with the likely result of i t s inter-turbine temperature (ITT) limits being exceeded.

Two system tests, a static and a dynamic system test, ensure system serviceability before take-off- The static test is performed with the engines running at idle, using a selector switch in the flight compartment. The dynamic test is conducted automatically by the APR controller, with the system armed and when the engines are accelerated through 83.5% Ni for take-off. The dynamic system test cannot be repeated unless the weight-on-wheels status changes or the system is selected off and re-armed.

4. ENGINE OIL SYSTEM (Figures 5 and 6)

Each engine is lubricated and cooled by its own self-contained oil system. In addition to an engine-mounted oil storage tank, an oil replenishment system located in the rear equipment bay is also provided.

Sensors are installed at suitable locations in the system to provide inputs to the oil temperature and pressure indicators in the flight compartment. Impending oil filter blockage is also monitored.

The principal components that make up the engine oil system are described in a flow sequence, starting at the oil replenishment system through to the oil return to the oil storage tank.

A. Oil Replenishment System

This system is used to add oil to the oil storage tanks on both engines. It consists of a replenishment tank, an electric pump, an OIL LEVEL CONTROL panel and a selector valve.

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OPERATING MANUAL PS? 601A-6

The OIL LEVEL CONTROL panel indicates if the storage tanks are full and also tests the replenishment system. The selector valve is manually selected to the desired left or right storage tank. The electric pump transfers oil from the replenishment tank to the selected storage tank.

The replenishment tank is gravity-filled. Its tank-mounted sight gauge indicates oil level.

B. Oil Storage Tank

Each storage tank oil level can be determined by a dips tic mounted in the filler cap. The tank can be directly gravity filled or remotely filled using the replenishment system.

C. Oil Temperature Sensor

This sensor, located in the storage tank and connected to an indicator in the f l i g h t compartment, allows monitoring of engine oil temperature including the oil cooler operation.

D. Oil Circulation

Oil flows from the storage tank to the lube pump. The pressurized oi l is then directed through the f i l t e r , the cooler and then to the various engine components requiring lubrication and cooling. Oil is returned to the storage tank by scavenge pumps.

E. Oil F i l t e r

The o i l f i l t e r consists of a f i l t e r element, a dif ferential pressure switch connected to an o i l pressure impending bypass indicator, and a bypass valve. The impending bypass indicator, located in the rear equipment bay, warns of an impending blockage of the f i l t e r element. Should the f i l t e r become clogged, the bypass valve would open to allow unfi l tered oi l to maintain engine lubrication.

F. Oil Cooler

The o i l cooler is an o i l - to - fue l heat exchanger which uses fuel as a cooling medium for the engine o i l .

G* Oil Pressure Sensor and Low Oil Pressure Switch

The pressure sensor and the low pressure switch, connected to their indicator and warning l ight respectively in the f l ight compartment, provide independent indications of engine oi l pressure. Both c i rcui ts sense the d i f ferent ia l oi l pressure between the lube pump discharge and the scavenge pump suction.

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ENGINE CONTROLS (Figure 7)

Each thrott le lever with i ts hinged thrust reverse (TR) lever is connected to i t s engine fuel control unit (FCU) through a single f lex ib le cable extending from the throttle quadrant in the f l ight compartment to the thrott le control gearbox on the engine.

This system transfers a l l thrott le and thrust reverse lever movement to the engine to command forward or reverse thrust as well as fuel shutoff. This system also mechanically provides a t a c t i l e feedback to the thrott le and thrust reverse levers when the FCU is kept a t , or should be returned to , idle by the t h r o t t l e retarder system.

Individual thrott le lever release latches prevent inadvertent selection of fuel SHUT OFF or fuel-on (IDLE) and individual thrust reverse lever release latches prevent inadvertent operation of thrust reverse levers. Mechanical interlocks within the thrott le quadrant also prevent a thrust reverse lever from being operated unless i ts thrott le lever is at IDLE, or prevent a throt t le lever from being advanced above IDLE when i ts thrust reverse lever is pulled up from the forward idle position.

A t h r o t t l e lever f r ic t ion adjustment control is also provided.

THRUST REVERSER (Figures 7 and 8)

Each engine is equipped with a thrust reverser to assist in a i rcraf t braking a f t e r landing. When the thrust reverser is deployed, a translating cowl moves rearward on tracks driven by a pneumatic actuator and uncovers forward facing cascade vanes. Blocker doors, operated by interconnected linkages, move inward to block the fan a i r exhaust duct and redirect the fan exhaust air through the cascade vanes. The system is powered by 14th stage bleed a i r .

A. Operation

Each thrust reverser is armed for operation by i ts associated REVERSE THRUST switch/light. With the throt t le lever at IDLE and either a weight-on-wheels signal or a 16-knot wheel spin-up signal, raising the thrust reverse lever to the deploy position in i t ia tes the following deployment sequence:

- Wing and engine anti- icing shutoff valves, i f open, are closed to conserve 14th stage bleed air for the reverser operation.

- 14th stage bleed air unlocks the reverser from i t s stowed position and powers a pneumatic drive unit (PDU) which deploys the reverser.

• As the reverser is fu l l y deployed, i t is locked in position by the PDU, and the thrust reverse lever is mechanically released by a throt t le retarder system and is e lect r ica l ly released by a throt t le solenoid, so that thrust settings up to maximum reverse thrust can be selected.

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Moving the thrust reverse lever back to the deploy pos i t i on t h r o t t l e s the engine down to reverse i d l e - Continued movement o f the lever t o the f u l l y down posi t ion stows and locks the reverser, i n the reverse sequence to the deployment described above.

As the reverser is f u l l y stowed and locked in pos i t i on by the PDU, the t h r o t t l e retarder system mechanically releases the t h r o t t l e lever so that forward thrust set t ing above IDLE can be selected again. Wing and engine an t i - i c i ng is also restored to normal cont ro ls .

After the above sequence to stow and lock the reverser i s complete, the reverser can be disarmed by i t s associated REVERSE THRUST sw i t ch / l i gh t .

B. Safety Features

Each reverser i s protected by the safety features described below.

(1) Throttle Retarder System

The th ro t t l e re tarder system is a mechanical system connected to the thrus t reverser and the engine FCU control l i nkage. I f the engine i s t h ro t t l ed above i d l e and there is an inadvertent th rus t reverser deployment, the t h r o t t l e retarder system returns the FCU to i d l e , and, through the i n t e r l i n k cable, pu l ls the t h r o t t l e lever to IDLE. This system also operates during normal operation of the reverser, as described in paragraph 6.A. above.

(2) Thro t t le Lockout System

The throttle lockout system prevents movement above IDLE of a previously retarded throttle lever if the aircraft is airborne and the thrust reverser moves from the fully stowed position. Under such conditions, initial movement of the thrust reverser energizes a throttle lockout solenoid which locks the throttle linkage at the FCU and prevents throttle lever movement beyond IDLE. If the reverser is returned to the stowed position, the throttle lockout solenoid is de-energized and freedom of movement is returned to the throttle lever.

(3) Auto Stow System

I n the case of an uncommanded movement of the reverser from the stowed pos i t i on , a microswitch commands the PDU to re tu rn the reverser to the stowed pos i t ion .

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Apr 02/87


(4) Emergency Stow System

If the REVERSER UNLOCKED light is on and the auto stow system fails to stow the reverser, the appropriate THRUST REVERSER EMERG STOW switch/light can be pressed to ensure positive operation of the sys^--to the stowed position.

(5) Safety Relay

Actuation of the thrust reverse lever with the aircraft not on the ground causes the REVERSE THRUST UNSAFE TO ARM light to come on.

ENGINE INSTRUMENTS (Figure 9)

Engine instruments monitor Nl %rpm, inter-turbine temperature (ITT), N2 %rpm, fuel flow, oil temperature and oil pressure.

A. Signal Data Converter (SDC)

The SDC controls the power supply and provides automatic dimming to the engine indicator systems.. Two power supplies are divided within the SDC into dual lamp-processing and signal-processing power supplies.

B. Engine Instruments

Each instrument provides a vertical analog display of the relevant engine variable using a series of miniature incandescent lamps inside the indicator, which provide the light source for the vertical row of coloured l ight segments.

Digital displays on the Nl, ITT, N2 and FUEL FLOW indicators provide more accurate indications when compared with the readings on the vertical scales.

ENGINE BLEED AIR

Engine bleed air consists of two systems, each with its own source. One source is at the 10th stage and the other is at the 14th stage of the compressor of each engine. Each system contains distribution ducting, shutoff valves, isolator valves and check valves. Both systems are protected by a bleed air leak detection system. The flight compartment BLEED AIR and ANTI-ICE control panels provide the necessary controls and indicators.

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canaaair ctiauencjer OPERATING KANUAL

PSP 601A-6

Tenth Stage Bleed Ai r (Figures 10 and 11)

The 10th stage bleed a i r system can be supplied from the l e f t and r i g h t engines or from the APU, or from a ground a i r supply un i t through an external connection on the lower l e f t side of the rear fuselage. The 10th stage system supplies bleed a i r to the fo l lowing systems:

A i r condi t ion ing/pressur izat ion

Cabin pressur izat ion control

Footwarmer/demister and emergency pressurization

Engine s t a r t i n g

A bleed a i r i s o l a t o r valve is normally closed to separate the l e f t and r i gh t d i s t r i b u t i o n duct ing. This iso lator valve is automat ica l ly opened by the engine s t a r t system to ensure a i r supply to both engines regardless of the a i r source. I t can also be selected open when requi red, i . e . , to supply both ACUs from a s ingle engine bleed source.

A l e f t and r i g h t pressure indicator receives signals from two sensors, ont on each side o f the i so l a to r valve, for continuous monitor ing. For operation of LH and RH footwarmer/demister valves and l e f t and r i g h t ACU valves, re fer t o Section 2.

Fourteenth Stage Bleed A i r (Figures 11 and 12)

The 14th stage bleed a i r system is supplied only by the l e f t and r i g h t engines. The 14th stage system supplies bleed a i r to the fo l low ing systems:

Wing a n t i - i c i n g

Engine a n t i - i c i n g

Thrust reverser

The operation o f engine and wing an t i - i c i ng valves ( inc lud ing the i so la to r valve) is con t ro l led by the ANTI-ICE control panel ( re fe r to Section 14).

Bleed A i r Leak Detection and Warning System (Figure 13)

Six temperature sensors ( f i r e - w i r e type) are attached to the bleed a i r ducts and are connected to two bleed a i r leak detection contro l u n i t s . Dual detect ion loops are provided for the l e f t and r i g h t sections of the 10th stage bleed a i r system, and single loops are provided f o r the 14th stage bleed a i r system, the fuselage pylons and the a n t i - i c i n g ducts running through the fuselage and wings. I f a leak occurs, the hot bleed a i r escaping i s detected by the temperature sensors and i n i t i a t e s a warning s igna l .

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Apr 02/87

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The warning signal is picked up by its leak detection control unit and transmitted to the flight compartment via a centrally located flashing DUCT FAIL warning light as well as an appropriate individual DUCT FAIL warning l ight on the control panel for the system affected by the leak.

The individual DUCT FAIL warning light identifies the defective duct whicn can then be depressurized and isolated using the BLEED AIR control panel.

A bleed air leak annunciator panel, behind the copilot's seat, also provides for fault isolation through eight latching magnetic indicators.

With the exception of the indicators on the bleed air leak annunciator panel, all of the warning indicators go out when their associated temperature sensors have cooled sufficiently. Warnings and testing of the bleed air leak detection system are summarized in Figure 13.

9. ENGINE STARTING AND IGNITION SYSTEMS (Figure 14)

The engine starting and ignition systems consist of a pneumatically driven air turbine starter, ignition-exciter boxes and igniter plugs for each engine. The systems are controlled by individual switch/lights in the flight compartment.

The starter transmits starting torque to the N2 rotor through the accessory gearbox. An automatic centrifugal shutoff switch opens at a preset rpm to protect the starter against overspeed. The 10th stage bleed air system supplies the starter through a starter valve.

Each engine has two ignition systems, A and B, each system connected to i ts igniter plug in the combustion chamber. The systems are powered by 115-volt ac power.

A. Ground Starting

The APU, an external ground air source or an operating engine can be used to supply the 10th stage bleed air system for engine starting.

An IGN A and/or IGN B switch/light arm(s) the associated igniter plugs on both engines. With the 10th stage bleed air system pressurized, pressing a START switch/light initiates the starting sequence on the associated engine. Refer to Figure 14 for description of the starting sequence.

B. In-Flight Starts

An IN FLIGHT START switch/light provides a separate power supply and fires both igniter plugs (A and B) on the selected engine without any other switch/light having to be operated- However, i f the windmilling rpm is less than 131 N2, starter assist using the START switch/light is required.

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Apr 02/87


C. Continuous Ign i t i on

A CONT IGN swi tch / l igh t provides continuous i gn i t i on to both engines through the pre-selected A and/or B i gn i t e r plugs. F i r ing o f the i gn i t e r plugs is continuous un t i l the CONT IGN swi tch / l igh t is pressed out .

10. ENGINE VIBRATION MONITORING SYSTEM (Figure 15)

The engine v ibra t ion monitoring (EVM) system provides a continuous ind ica t ion of the v ib ra t ion level of each engine. The main components of the system include a transducer mounted on the compressor casing of each engine, a signal condit ioner and an indicator panel in the f l i g h t compartment.

Each transducer generates an e lec t r i ca l signal proport ional t o the i n tens i t y o f engine v ib ra t i on . The signal condit ioner converts these signals i n t o values readable on the EVM ind ica tor .

An alarm c i r c u i t causes an amber caution l i g h t on the EVM ind i ca to r panel to come on i f the v ib ra t ion level of e i ther engine exceeds 1.7 MILS f o r a period greater than 3 seconds. This 3-second delay prevents spurious warnings caused by high t ransient engine v ibra t ions.

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Apr 02/87


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Power Plant - Schematic Figure 1

SECTION 17 Page 12

Apr 02/87


BLEED AIR f

4-

ECOLOGICAL DRAIN TANK

FIREWALL! SHUTOFF

(VALVE

AIRCRAFT FUEL SYSTEM

W

1 1 Hftl I j

F™5H

TO RIGHT ENGINE

F̂ WBT

e=FECTIVnY: A C 6001TO 613*

LEFT ENGINE ILLUSTRATED

LEGEND • I • I • MOTIVE FUEL SUPPLY

i l l l l l l l f l t l METERED PRESSURIZED FUEL

* * * * * * ECOLOGICAL DRAIN LINE

ftT] DIFFERENTIAL PRESSURE SWITCH

Engine Fuel System - Schematic Figure 2 (Sheet 1)

SECTION 17 Page 13

Apr 10/95


1 j MAI«=OU)~H

UlLUHVUV; A/C619S AND SUBS

LEFT ENGINE ILLUSTRATED

LEGEND • ( • I B MOTIVE FUEL SUPPLY millttllir METBIED PRESSURIZED FUEL

fp"] OfFFERBCTIAL PRESSURE SWITCH

Engine Fuel System - Schematic Figure 2 (Sheet 2)

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Apr 10/95


FUELTEMPERATURE INDICATOR

Shows temperature at left and right toef heater outlets.

A!C 6001 TO 5134: Nofnw opefflDHQ rai Cwbcmry range ftraDoer} and

>VC 5136 AND 8UB8:

r e to acre -2CrCto5*C 6CTCto70*C

No*nrfop*rata^ range (ferem) 4*Cto120*C Ceitfomry range fj^iiow) -66~Cto4*C

FUEL CONTROL PUSH ON/OFF 1

RTANK

PUMP EJCTRS

d

LOW PRESSURE WARNING UGHTS

Amber warning light comes on to indicate low pressure at associated engine fuel inlet port.

VALVE CLOSED UGHTS

White light comes on whenever associated firewall fuel ahutoff valve is closed.

FILTER BYPASS WARNING UGHTS

Amber light comes on when fuel pressure drop is detected across associated main fuel fitter.

NOTES

1 Refer to FUEL for details of aircraft fuel system control and rnonrtoring.

2 On A/C 5135 and SUBS, the fuel LOW PRESS lights wiH be on until the pumps are selected ON.

CENTRE INSTRUMENT PANEL

Fuel Control Panel -Engine Fuel System Monitoring

Figure 3 SECTION 17

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OPBUTMG MANUAL PSP 601A-6

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SECTION 17 Page 14B

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READY LIGHT

Green light comes on to confirm system readiness after system is initially armed. When APR operates, light goes out when L ON or R ON light comes on.

L ON/R ON LIGHTS

Green lights come on to indicate left or right engine is responding to APR command following a power loss.

APR SELECTOR SWITCH

Three-position toggle switch:

ARM - Arms system if both ENG SPEED CONTROLS switches are on, both engines are in N l speed control mode and APR light is out.

OFF - De-activates system.

TEST/RESET - Initiates static test of system (refer to NORMAL PROCEDURES). Resets system after a fault is cleared.

APR LIGHTS

Amber light comes on when:

- APR selector switch is off prior to take-off.

- APR self-monitoring circuits detect fault in APR or engine fuel control system.

READY

li TEST I

] I APR

ENG. SPEED CONTROL ON ON

Green light comes on during static and dynamic tests of the system (refer to NORMAL PROCEDURES).

ENGINE SPEED CONTROL SWITCHES

Two-position toggle switches.

ON - Engine speed control i s m N l mode when NT rpm exceeds nominal 79.1 %.

OFF - Engine speed control is in N2 mode regardless of N l rpm.

CENTRE I N S T R U M E N T PANEL

APR and Engine Speed Control Panel SECTION 17 Figure 4 Page 15

Apr 02/87


t

LD BYPASS VALVE

DIFFERENTIAL PRESSURE SENSOR OR SWITCH

Engine Oil System - Schematic Figure 5

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Apr 02/87


OIL PRESSURE INDICATOR

Vertical scale indicator displays ofl pressure of each engine. Coloured fight segments of vertical scales come on to indicate the following range.

NZ 5001 TO 5134: Low pressure warning Ene (red) 25 psi Normal operating range (green) 25 to 95 psi Cautionary pressure range (yellow) 95 to 100 psi High pressure warning line (red) 100 psi

OIL TEMP

—»1t0 — 1—140 — |«» «» IOMI mmm \mmm ^m

— 120 — «^» «a» ! « • • SMS «•» mm — 100«-» !•» ^» »••• GO sea • • • OBi aa» 0 aaa aaaB -70 - I B*

0 jd j

j 1 q 1 IJ 1 •i I

J 1

"1 1 R

A/C 5135 AND SUBS:

Low pressure warning fine (red) 25 psi Normal operating range (green) 25 to 115 psi Cautionary pressure range (yellow) 115 to 130 psi High pressure warning line (red) 130 psi

OIL TEMPERATURE aMDICATOft

Vertical scale M o t o r dspisyt oil Coloured fight segments of vertical come on to indicate the

LOW OIL PRESSURE LIGHTS

Red warning lights coma on whan oil pressure of associated angina drops below 2B i . 3 pax.

Normal operaUiy lenye tarean) ^M6onajYranga (yaaow) Warning fine (red)

-20°Cto140°C 140°Cto180°C

ieo°c

ENGINE INSTRUMENT PANEL

Oil Temperature and Pressure Indicators SECTION 17 Figure 6 p a g e 17

Apr 10/95


THRUST REVERSER EMERG STOW SWITCH/LIGHTS

When pressed, power is applied directly to arming and stow solenoid valves to initiate stowage of reverser.

Amber REVERSE UNLOCKED light comes on whenever reverser moves from fully stowed position and remains on until reverser is returned to fully stowed position.

Green REVERSE THRUST light comes on when reverser reaches fully deployed position and goes out immediately when reverser moves from deployed position.

THRUST REVERSER EMERG STOW

REVERSER UNLOCKED!

REVERSE THRUST

REVERSER UNLOCKED

REVERSE THRUST

THRUST REVSISE (TR) LEVERS

With throttle levels at IDLE, putting onTR levels deploy levsisers if following conditions met:

- REVERSE THRUST switch/lights have armed reversers.

• Aircraft on ground or wheel spin-up exceeds 16 knots.

Throttle solenoids prevent TR lever movernem beyond deploy (or reverse idle) position until reverser assemblies fully deployed.

Once reversers fully deployed, TR levers regulate reverse thrust from reverse idle to maximum reverse power.

Reverser operation shuts off 14th stage bleed air to engine and wing

GO-AROUND SWITCHES

Momentarily push button switches. These switches are associated with go-around mode of flight director system.

Returning TR levers to forward IDLE (fully down) stow reversers. Once reversers stowed, throttle levers can be moved forward to increase thrust.

NOTE Reverser deployment do not prevent throttle levers from being selected to SHUTOFF.

THROTTLE SETTINGS

SHUTOFF - Shuts off fuel to engine at the FCU. Located at rear throttle

IDLE - Lowest forward thrust setting. Located at idle throttle lever stop.

MAX POWER - Highest forward thrust setting. Located at forward throttle lever stop.

PUSH LEFT PUSH RIGHT

GLARESHIELD

THROTTLE LEVERS

Control forward thrust and acts a fuel shutoffs. Remain locked at IDLE position during thrust reverser operation.

THROTTLE LEVER RELEASE LATCHES

Lift to advance throttle levers from SHUTOFF to IDLE positions or retard throttle levers from IDLE to SHUTOFF positions.

THRUST REVERSE LEVER RELEASE LATCHES

Lift to release TR levers from forward IDLE stops.

^ 3 THROTTLE LEVER FRICTION ADJUSTMENT

Adjusts friction on throttle levers only. Rotate control clockwise to increase friction.

CENTRE PEDESTAL

REVERSE THRUST SWITCH/UGHTS

When pressed in, arms thrust reverser system and puts on amber ARMED light.

When pressed out, providing thrust reverser stowed* disarms thrust reverser system and puts out amber ARMED tight.

Amber UNSAFE TO ARM light comes on if:

- electrical fault exists in reverser

F tEVERSE THRUST 3\ LEFT RIGHT ( j

| UNSAFE 1 j TO ARM j

ARMED

1 PUSH TO

1 UNSAFE 1 j TO ARM j

j ARMED

ARM

deploy is selected or deploy switch fault occurs during flight.

CENTRE PEDESTAL

Throttle Quadrant and Thrust Reverser Controls and Indicators

Figure 7

SECTION 17 Page 18

Apr 02/87

canatiair ctiauenejer


STOWED POSITION CASCADE VANES

FAN AIR

BLOCKER DOORS

TRANSLATING COWL

TORQUE BOX

DEPLOYED POSITION

Thrust Reverser Stowed and Deployed Positions SECTION 17 Figure 8 Page 19

Apr 02/87


Engine Instruments and Control Panel Figure 9

SECTION 17 Page 20

Apr 10/95


GROUND AIR SUPPLY

A TO CABIN PRESSURIZATiON ? CONTROL SYSTEM

| JET PUMP

LEFT ENGINE 10TH STAGE

BLEEDS

RIGHT ENGINE 10TH STAGE

BLEEDS

ENGINE START

SYSTai

CLOSE

EFFECTIVITY

H A/C 5001 TO 6134

LEGEND

BLEED AIR

ELECTRICAL SIGNAL

SHUTOFF VALVE

REGULAT1NG/SHUTOFF VALVE

CHECK VALVE

Tenth Stage Engine Bleed Air - Schematic Figure 10

SECTION 17 Page 21

Apr 10/95

OPERATING MANUAL PSP 6 0 1 A - 6

10TH AND 14TH STAGE SWITCH/UGHTS

I in, associated blood air shtitoff valve opens and vriata BLEED CLOSED tight goes out Whan pressed out valve

tand light comas on.

BLEED AIR ISOL SWITCH/UGHT

Whan prassad in, bleed air isolator vafrve opens. When prassad o u t valve closes.

Green OPEN light comes on whenever bleed air aoiator valve is open.

BLEED AIR PRESSURE GAUGE

Indicates pressure in left and right sections of 10TH stage bleed »r system.

EFFECTMTY: A/C 5135* SUBS

OVERHEAD PANEL

Bleed Air Control Panel SECTION 17 Figure 11 Page 22

Apr 10/95


TOWING ANTMCING

TO ENGINE ANTI-ICING

TO THRUST REVERSER

LEFM4TH STAGE BJGINE BLSD

TOWING ANTMCING

TO ENGINE ANTI-ICING

TO THRUST REVERSER

CLOSE -**

[ CLOSE THRUST REVERSERS

IN OPERATION

CLOSE

RIGHT 14TH STAGE ENGINE BLEED

O

CLOSE

U-+ RH ENG RRE

PUSH

LEGEND

BLEED AIR

ELECTRICAL SIGNAL

SHUTOFF VALVE

REGULATING/SHUTOFF VALVE

CHECK VALVE

Fourteenth Stage Engine Bleed Air - Schematic Figure 12

SECTION 17 Page 23

Apr 10/95


DUCT MON SWTTCH

Three-portion DUCT MON toggJe switch tests serviceability of each of the detector loops A and B on the left and right 10th stage manifold sections.

LOOP A - Duct fail warning occurs if loop A of either section is damaged.

LOOP B - Duct fail warning occurs if loop B of either section is damaged.

BOTH - In-flight switch position. Both detection loops are in operation on left and right sections.

DUCT FAIL LIGHTS (4)

Red light comes on if the bleed leak temperature sensors detect a failure in the associated duct segment.

Light goes out when the failed duct is isolated and temperature sensor cools.

OVERHEAD PANEL

WING ANTI-ICE DUCT FAIL LIGHT

Red DUCT FAIL light comes on if bleed air leak is detected in wing left and right anti-icing ducts running along fuselage.

OVERHEAD PANEL

, 14 STAGE , - , * RIGHT LEFT '

o o • RIGHT F U S LEFT

o o (RIGHT

O 'RIGHT

o

- W I N G *

-10 STAGE

IND RESET SYSTEM TEST

O O BLEED AIR LEAK

8LEED AIR LEAK DETECT

DUCT

FAIL

PUSH TO TEST

BLSD AIR LEAX DETECT SWITCH/LIGHT

Red DUCT FAIL light flashes if a bleed air leak is detected by any of the detection elements.

PUSH TO TEST-- When pressed, system is tested by grounding detection circuit to simulate bleed air leak. Flashing DUCT FAIL light on swttch/tight and steady DUCT FAIL lights on bleed air and arm-ice panels come on if leak detection system is serviceable.

BLEED AIR LEAK ANNUNCIATOR PANEL

Panel indicate* s have two positions: a black set position when no fault exists and a white reset position visible when there is a bleed leak in the associated ducting.

Reset positions VB magneocaUy latched to remain on after associated temperature sensor has cooled or electrical power is removed from aircraft. Pressing IND RESET button returns positions to set.

Pressing SYSTfcM TEST switch tests system by grounding detection circuit to simulate bleed air leak. Afi the DUCT FAIL lights come on and ail eight indicatots on panel show white if leak detection system is serviceable.

CENTRE INSTRUMENT PANEL

Bleed Air Leak Warning and Testing SECTION 17 Figure 13 Page 24

Apr 02/87

cttanenejer OPERATING MANUAL

PSP 601A-6

IGNITION SWITCH/LIGHTS

When pressed in, arms associated igniter plug of both engines for start and continuous igntion operation.

When pressed out, disarms associated igniter plug of both engines for start and continuous ignition operation.

Green IGN A (or SGN B) light comes on immediate associated switch/light pressed.

White ON lightfs) comets) on when associated igniter plugs on one or both of the engines are in operation.

START SWITCH/UGHTS

Momentarily pressing switch/light causes green START light to come on and initiates engine start sequence:

- opens Of not previously opened) left and right bleed air shutoff valves and isolator valve.

- opens associated starter valve.

- fires pre selected A and/or B igniter plugis) on associated engine, ignition white ON tightts) cornels) on.

When engine reaches 55° N2, the starter automatic shutoff switch:

- closes (unless selected open by associated control switch) left and right bleed air shutoff valves and isolator valve.

doses associated starter valve.

- turns off pre selected A and/or B igniter plugis) of 1 engine, ignition white ON bghtls) goies) out.

STOP SWITCH/LIGHTS

Momentarily pressing switch/light stops engine start

CONT IGN SWITCH/LIGHT

When pressed in. green CONT IGN light comes on and continuous ignition is supplied to both engines through IGN A and/or IGN B switch/tight(s). When pressed out. CONT IGN light goes out and continuous ignition is turned off.

IN FUGHT START SWITCH/UGHTS

When pressed in, fires both igniter plugs on associated engine and green IN FUGHT START light and white ON light come on.

^A/hen pressed out, turns off both associated igniter plugs, green IN FUGHT START light and white ON light.

Amber STOP light comes on GO seconds after START switch is pressed if engine has failed to start.

OVERHEAD PANEL

Engine Start and Ignition Controls SECTION 17 Figure 14 Page 25

Apr 02/87


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ENGINE VIBRATION

OEEN

EFFECTTVTTY: A/C 5001 TO €104

ENGINE VIBRATION CAUTION UGHT

Amber HIGH VIS fight indicates that vibration leva* of one or both of the engines hat exceeded 2.7 MILS (FAN) cr 1.7 IfiLS (CORE) *x3eeoond».

P i w i n j HIGH VIB indicator switch provides functional test of sidicatof and signal conditioner. Indicator VMU show L and R reading of &S ± OJg MILS and, if switch is HI ssssrt for 3 1 HIGH VIB fight comas on.

ENGINE VIBRATION INDICATOR

Provides continuous mdteetion of left and fight J by transducers mounted on engine

engine vibration

NOTE

Press to test function transducers or cable must be checked by

not verify operation of Operation of these

indicator reading after

engine mounted

B7ECTMTY: A/C 5105 AND SUBS

1.7 MILS. Amber fan

I H B I G HI V © AND RH B I G W VIB SVVITCH/UGHTS

AinberoBfB) aghtl Off •Ontl

12.7MBJB. PieesliHjCOriEFANI

lnoloatDrwMarKiwLandRreoclngof3^ta2MUJS and.lfaeAchlspteeeedsySsecondsCOREFAN Ight comas on. PILOTS SIDE PANEL

Engine Vibration Monitor Panel SECTION 17 Figure 15 Page 26

Apr 10/95

UTHO CANADA

Documents

5. ENGINE CONTROLS 6 - · PDF fileThe flight compartment throttle lever movement is transmitted to the FCU which in turn controls engine speed in one of the following two modes: