De Havilland DASH8 100 Series Maintenance Manual

DSERIES 100

AIRCRAFT MAINTENANCE MANUAL

**ON A/C ALL

ENGINE FUEL INDICATING − GENERAL

1. General Engine fuel indicating associated with the engine fuel and control system consists of a fuel

low pressure warning system, a fuel flow indicating system and a fuel temperature indicating system for each engine. The systems for each engine have like components and are identical in operation.

Fuel low pressure warning consists of a fuel low pressure switch, sensing fuel pump inlet pressure and connected by electrical wiring to the caution lights system. Fuel flow indicating is by means of a transmitter sensing fuel flow to the engine fuel manifold, connected by electrical wiring to an indicator on the engine instrument panel. Fuel temperature indicating is by means of a transmitter sensing temperature at the fuel pump inlet, connected by electrical wiring to an indicator on the engine instrument panel.

For location of the fuel low pressure switches, and fuel flow and fuel temperature transmitter, refer to ENGINE FUEL AND CONTROL − GENERAL, Figure 1.

Other indications associated with the fuel system are fuel quantity indicating systems, one for each fuel tank; refer to Chapter 28.

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FUEL LOW PRESSURE WARNING − DESCRIPTION AND OPERATION

1. General Two identical but separate fuel low pressure warning systems, monitor input fuel pressure to

the engine driven fuel pump for each engine. In the event fuel pressure falls below a pre−set minimum pressure, the system activates caution lights.

2. Description Refer to Figure 1. The fuel low pressure warning is provided by a fuel low pressure switch. The switch is

installed in a tapping in the fuel/filter housing, upstream of the engine driven fuel pump. Each switch is connected to the appropriate ENG FUEL PRESS negative seeking caution

light circuit in the caution lights panel in the flight compartment. For details of the caution lights system, refer to Chapter 33.

3. Operation Refer to Figure 2. The switch is provided with one contact set. The switch mechanism is calibrated to operate

the contact from C to N.O. on increasing fuel delivery pressure of 7.5 + or − 0.8 psig and from C to N.C. at 5.5 + or − 0.8 psig decreasing pressure.

The N.C. contact is connected to ground and the N.O. contact is an open circuit. With fuel delivery pressure above 7.5 psig, the ground is removed and the associated caution

light goes out. Should the fuel delivery pressure fall below 5.5 psig, the caution light logic circuit is connected

to ground and the caution light comes on.

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FUEL LOW PRESSURE WARNING

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Figure 1

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MASTER EFFECTIVITY: See Pageblock 73−31−00 page 1

FUEL LOW PRESSURE WARNING

Figure 2

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FUEL FLOW INDICATING − DESCRIPTION AND OPERATION

1. General Fuel flow indicating for the engine fuel and control installation consists of two identical, but

independent fuel flow indicating systems, one for each engine. Each system consists of a fuel flow transmitter connected to a fuel flow indicator. The transmitters are mounted on the engine casings and the indicators are located on the engine instrument panel in the flight compartment.

The electrical circuit for each engine is powered from the 28 V dc bus, and protected by a 5−ampere FUEL FLOW IND ENG 1 (C5) left main bus or FUEL FLOW IND ENG 2 (Q5) right main bus circuit breaker.

2. Description and Operation Refer to Figure 1. A. Fuel Flow Transmitter

A fuel flow transmitter installation consisting of a transmitter and a mounting bracket assembly is attached to the engine front inlet case rear flange at approximately 3 o’clock (viewed from rear of engine). The transmitter is connected into engine metered fuel lines between the hydromechanical fuel control unit (HMU) and the fuel manifold and nozzles installation of each engine fuel system.

A directional arrow, on the transmitter casing, points away from the HMU and provides indication of required fuel flow direction. The bracket assembly and the transmitter are provided with an uneven bolting pattern to prevent incorrect installation.

The transmitter is of the rotating impeller type and is connected to the associated indicator by electrical wiring.

B. Fuel Flow Indicators

Refer to Figure 2 and The fuel flow indicators are located on the engine instrument panel in the flight

compartment and are internally lighted by white lighting, powered from the 5 V dc lighting system (refer to Chapter 33).

Each indicator receives an ac signal equivalent to fuel flow from its associated transmitter and provides a fuel flow readout by means of a moving pointer against a fixed dial and an equivalent digital display. A 0 to 5 V dc signal proportional to fuel flow is relayed to the flight data recorder (refer to Chapter 31).

The digital display is of the 4−digit liquid crystal display (LCD) type. Pre mod 8−1/1101 aircraft incorporate a heater associated with LCD operation in each indicator. The heaters are powered from the 28 V dc secondary buses, engine No. 1 IND LCD HTR (S2) indicator heater from the left bus and engine No. 2 IND LCD HTR (A1) indicator heater from the right bus (refer to Chapter 77). The heaters are temperature controlled to ensure LCD operation at low temperature conditions in the flight compartment.

The indicator dial is marked FF PPH × 100 (KG/H × 100) major graduations in increments of 100 PPH (100 KG/H) between 0 and 1200 PPH (0 and 550 KG/H), with minor graduations in increments of 50 PPH (50 KG/H). Indicator for engine No. 1 is powered by 28V dc L MAIN BUS (C5). Engine No. 2 by 28V dc R MAIN BUS (Q5).

Figure 3.

and the scale is graduated in

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A test pushbutton is provided on the indicator, which when pressed, with power on, causes the indicator to move to 1050 PPH (400 ±70 KG/H), indicated by a blue dot on the face of the indicator dial and an equivalent digital display. Releasing the pushbutton causes the pointer to fall to zero and a digital display of 0.0.

Loss of electrical power to the indicator causes the pointer to move off scale to below zero and the digital display is blanked.

Mod 8/1738 introduces fuel flow indicators with specification changes as follows:

(1) FF (fuel flow) of 1200 PPH (550 KG/H)

read 1200 PPH (550 KG/H). (2) Low range indication is extended from 80 PPH (35 KG/H) to 40 PPH (20 KG/H). (3) The FDR outputs at pins J and K shall be 5V dc (max) for fuel flows of 1200 PPH

(550 KG/H) and higher. (4) Normal fuel flow indications will be available without delay when fuel flow is reduced

below 1200 PPH (550 KG/H)

or less.

and higher (beyond indicator range) shall

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FUEL FLOW TRANSMITTER INSTALLATION

Figure 1

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FUEL FLOW (FF) INDICATING SYSTEM − SCHEMATIC

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Figure 2

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FUEL FLOW

KG/H x 100

INDICATOR FUEL FLOW (METRIC)

5V DC LTG

28V DC L BUS

2 5

NOTE

1. Engine No. 1 indicating system shown, engine No. 2 system similar.

2. Identification code is 7331.

(REFER TO CHAPTER 33)

FUEL FLOW IND ENGINE 1

0 TO 5V DC

OUTPUT TO FLIGHT DATA RECORDER

FUEL USED PULSE

DC REF

J K L M

A C B

S F

G H E

N P U

V R

FUEL FLOW INDICATOR

dam01_7332000_003.dg, gw, 31/10/01

FUEL FLOW (FF) INDICATING SYSTEM (METRIC) − SCHEMATIC

MASTER EFFECTIVITY: See Pageblock 73−32−00 page 1

Figure 3

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FUEL FLOW INDICATING − MAINTENANCE PRACTICES

1. General A. The maintenance procedures that follow are for the:

–

Removal and installation of the fuel flow indicator

–

Removal and installation of the fuel flow transmitter

–

Operational test of the fuel flow indicator.

2. Removal/Installation A. Removal of the Fuel Flow Indicator

Refer to Figure 201.

WARNING

(1) Remove the electrical power from the aircraft (refer to AMM 12−00−05). (2) Open, safety clip and tag the circuit breakers that follow:

(a) On the left DC circuit breaker panel:

(b) On the right DC circuit breaker panel:

(3) Remove the fuel flow indicator (1) as follows:

(a) Loosen the top right and lower left screws (3), that attach the indicator (1),

(b) Push in the top right and lower left screws (3) to disengage the clamping

(c) Pull out the fuel flow indicator (1) from the instrument panel (4). (d) Disconnect the electrical connector (2), and remove the indicator (1). (e) Put a protective cap on the electrical connector (2).

B. Installation of the Fuel Flow Indicator

Refer to Figure 201.

(1) Make sure that (2) Install the fuel flow indicator (1) as follows:

(a) Remove the protective cap from the electrical connector (2). Connect the

(b) Insert and align the indicator (1) on the instrument panel (4).

: REMOVE ALL ELECTRICAL POWER FROM THE AIRCRAFT

BEFORE YOU DO MAINTENANCE. PUT WARNING PLACARD AT THE EXTERNAL POWER RECEPTACLE AND IN FLIGHT COMPARTMENT. THIS IS NECESSARY TO PREVENT ELECTRICAL SHOCK TO PERSONS AND/OR DAMAGE TO THE EQUIPMENT.

–

FUEL FLOW IND ENG 1 (C5)

–

IND LCD HTR (S2)

–

FUEL FLOW IND ENG 2 (Q5)

–

IND LCD HTR (A1)

approximately five turns.

mechanism.

the aircraft is in the same configuration as in the removal procedure.

electrical connector (2) to the indicator (1).

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(3) Remove the safety clips and tags and close the circuit breakers that follow:

(a) On the left DC circuit breaker panel:

–

FUEL FLOW IND ENG 1 (C5)

–

IND LCD HTR (S2)

(b) On the right DC circuit breaker panel:

–

FUEL FLOW IND ENG 2 (Q5)

–

IND LCD HTR (A1) (4) Remove all the tools, equipment and unwanted materials from the work area. (5)

Do the operational test of the fuel flow indicator (refer to Para 3. A).

C. Removal of the Fuel Flow Transmitter

Refer to Figure 202 Equipment and Materials:

–

Fuel resistant container, commercially available

WARNING: REMOVE ALL ELECTRICAL POWER FROM THE AIRCRAFT

BEFORE YOU DO MAINTENANCE. PUT WARNING PLACARD AT THE EXTERNAL POWER RECEPTACLE AND IN FLIGHT COMPARTMENT. THIS IS NECESSARY TO PREVENT ELECTRICAL SHOCK TO PERSONS AND/OR DAMAGE TO THE

EQUIPMENT. (1) Remove the electrical power from the aircraft (refer to AMM 12−00−05). (2) For the No. 1 fuel flow transmitter, open the access panel 414AR to get access to

the transmitter (refer to AMM 06−40−10).

(3) For the No. 2 fuel flow transmitter, open the access panel 424AR to get access to

the transmitter (refer to AMM 06−40−10).

(4) Remove the fuel flow transmitter (1) as follows:

(a) Disconnect the electrical connector (2) from the fuel flow transmitter (1). Put a

protective cap on the electrical connector (2). (b) Put a container below the fuel flow transmitter (1) to collect the fuel that spills. (c) Disconnect the top fuel line (3) from the fuel flow transmitter (1). Install a

protective cap on the open end of the top fuel line (3). (d) Disconnect the bottom fuel line (4) from the fuel flow transmitter (1). Install a

protective cap on the open end of the bottom fuel line (4). (e) Hold the fuel flow transmitter (1) and remove the bolts (5) and washers (6) that

attach the fuel flow transmitter (1) to the bracket (7). (f) Remove the fuel flow transmitter (1). (g) Install protective caps on the electrical connector and the open ends of the fuel

flow transmitter (1).

D. Installation of the Fuel Flow Transmitter

Refer to Figure 202

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Equipment and Materials:

–

Fuel resistant container, commercially available (1) Make sure that the aircraft is in the same configuration as in the removal procedure. (2) Install the fuel flow transmitter (1) as follows:

(a) Put a container below the fuel flow lines (3) and (4) to collect the fuel that

spills.

(b) Remove the protective caps from the electrical connector and the open ends of

the fuel flow transmitter (1).

such that the flow arrow on the fuel flow transmitter (1) points down. Install the fuel flow transmitter (1) to the bracket (7) with the washers (6) and bolts (5).

(d) Torque the bolts (5) to 18 to 22 pound−inches (2.03 to 2.49 N·m) (refer to

AMM 20−14−01).

(e) Remove the protective cap from the bottom fuel line (4). Connect the bottom

fuel line (4) to the fuel flow transmitter (1).

(f) Remove the protective cap from the top fuel line (3). Connect the top fuel line

(3) to the fuel flow transmitter (1).

(g) Remove the protective cap from the electrical connector (2). Connect the

electrical connector (2) to the fuel flow transmitter (1). (3) Remove all the tools, equipment and unwanted materials from the work area. (4) Do the engine start procedure for the related engine (refer to AMM 71−00−00). (5) Examine all the connections at the fuel flow transmitter and the tube assemblies for

signs of fuel leaks. (6) Do the engine shutdown procedure for the related engine (refer to AMM 71−00−00). (7) For the No. 1 fuel flow transmitter, close the access panel 414AR (refer to AMM

06−40−10). (8) For the No. 2 fuel flow transmitter, close the access panel 424AR (refer to AMM

06−40−10).

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PPH x 100

1. Indicator.

2. Electrical connector.

3. Screw.

LEGEND

4. Instrument panel.

PPH x 100

dam01_7332004_002.dg, gw, aug14/2008

Fuel Flow Indicator − Removal and Installation

MASTER EFFECTIVITY: See Pageblock 73−32−00 page 201

Figure 201

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FRONT INLET CASE

REAR FLANGE (REF)

LEGEND

1. Fuel flow transmitter.

2. Electrical connector.

3. Fuel line.

4. Fuel line.

5. Bolt.

6. Washer.

7. Bracket.

dam01_7332002_001.dg, vk, jul24/2009

FUEL FLOW TRANSMITTER − REMOVAL/INSTALLATION

MASTER EFFECTIVITY: See Pageblock 73−32−00 page 201

Figure 202

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3. Adjustment/Test A. General

(1) The maintenance procedure that follows is to make sure that the fuel flow indicators

operate correctly. The test procedure includes the operational test of the fuel flow indicator.

B. Operational Test of the Fuel Flow Indicator

WARNING: MAKE SURE THAT ALL PERSONS IN THE AREA OF THE

AIRCRAFT ARE TOLD BEFORE THE ELECTRICAL SYSTEM IS ENERGIZED. IF YOU DO NOT DO THIS, YOU CAN CAUSE

INJURIES TO PERSONS OR DAMAGE TO EQUIPMENT. (1) Apply electrical power to the aircraft (refer to AMM 12−00−05). (2) Push and hold the test button on the indicator. Make sure that:

–

The pointer moves to the blue dot on the indicator face.

–

The digital display of the indicator reads 1050.

(3) Release the test button. Make sure that:

–

The pointer of the indicator moves to zero.

–

The digital display of the indicator reads 0000.

(4) Remove the electrical power from the aircraft (refer to AMM 12−00−05).

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FUEL TEMPERATURE INDICATING − DESCRIPTION AND OPERATION

1. General Refer to Figure 1. The fuel temperature indicating system for the fuel feed systems of both engines consists of

temperature transmitter for each engine connected by electrical wiring to a dual indicator which serves both engines.

2. Description and Operation A. Fuel Temperature Transmitter

A fuel temperature transmitter is located in an engine supplied tapping upstream of the engine driven pump (see ENGINE FUEL AND CONTROL − GENERAL, Figure 1). The transmitter is a standard electrical resistance type component, in which resistance varies in direct proportion to sensed temperature.

B. Fuel Temperature Dual Indicator

The fuel temperature indicator for both engines is mounted on the engine instrument panel in the flight compartment.

The indicator contains separate circuits and indicating mechanisms for engine No. 1 and engine No. 2, which receive separate inputs from the associated transmitter. The circuits are separately powered from the 28V dc buses and protected by 5−ampere FUEL TEMP ENG 1 (N4) left secondary bus and FUEL TEMP ENG 2 (E4) right secondary bus, circuit breakers. Lighting for the indicator is from the 5V dc lighting system (refer to Chapter

33). The indicator provides a readout of fuel temperature by means of moving pointers

against a fixed dial. The pointers for engines No. 1 and No. 2 are mounted on concentric shafts and identified 1 and 2 by a marking on the dial for each half scale.

The indicator dial is marked FUEL TEMP with the two separate half scales marked degrees C and graduated from −40 to +80 degrees, with major graduations increments of 40 degrees and minor graduations in increments of 10 degrees.

On Pre Mod 8/0861 aircraft the scale is provided with a green arc from +11 to +43 degrees with a red limit mark at 57 degrees, and yellow arcs +43 to +57 degrees and

−40 to +11 degrees. On Mod 8/0861 aircraft, the scale is provided with a green arc from +11 degrees to +57

degrees, yellow arc from −40 degrees to +11 degrees and red limit mark at 57 degrees. In the event of indicator power supply failure, the pointer(s) of the affected circuit(s)

move off scale below −40 degrees.

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FUEL TEMPERATURE INDICATING SCHEMATIC

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Figure 1

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FUEL TEMPERATURE INDICATING − MAINTENANCE PRACTICES

1. Removal/Installation The following maintenance procedures are approved for the removal and the installation of

the fuel temperature dual indicator. A. Removal of the Fuel Temperature Dual Indicator

Refer to Figure 201.

WARNING

(1) Remove electrical power from the aircraft. (2) Open, safety clip and tag the following circuit breakers:

: REMOVE ALL ELECTRICAL POWER FROM THE AIRCRAFT

BEFORE YOU DO MAINTENANCE. PUT WARNING PLACARD AT THE EXTERNAL POWER RECEPTACLE AND IN FLIGHT COMPARTMENT. THIS IS NECESSARY TO PREVENT ELECTRICAL SHOCK TO PERSONS AND/OR DAMAGE TO THE EQUIPMENT.

CIRCUIT BREAKER PANEL

LEFT DC N4 FUEL TEMP ENG 1

RIGHT DC E4 FUEL TEMP ENG 2

(3) Remove the fuel temperature indicator as follows:

(a) Loosen the indicator top right securing screw approximately five turns. (b) Push in the securing screws to disengage the clamping mechanism. (c) Slide out the fuel temperature indicator. (d) Disconnect the connector and remove the Fuel Temperature Indicator.

B. Installation of the Fuel Temperature Dual Indicator

Refer to Figure

(1) Install the fuel temperature indicator as follows:

(a) Make sure the aircraft is in the same configuration as in the removal

(b) Connect the connector to the fuel temperature indicator. (c) Insert and align the fuel temperature indicator. (d) Tighten the fuel temperature indicator top right securing screw.

(2) Remove the safety clips and tags and close the following circuit breakers:

CIRCUIT BREAKER PANEL

LEFT DC N4 FUEL TEMP ENG 1

RIGHT DC E4 FUEL TEMP ENG 2

201.

procedure.

CIRCUIT BREAKER NO. NAME

(3) Remove all tools, equipment and unwanted materials from the work area.

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WARNING: MAKE SURE THAT ALL PERSONS IN THE AREA OF THE

AIRCRAFT ARE TOLD BEFORE THE ELECTRICAL SYSTEM IS ENERGIZED. IF YOU DO NOT DO THIS, YOU CAN CAUSE

INJURIES TO PERSONS OR DAMAGE TO EQUIPMENT. (4) Apply electrical power to the aircraft. (5) If the fuel temperature indicator is unserviceable or there is no power supply to the

indicator, the pointer on the indicator will move off the scale below −40°C (−40°F).

(6) If the fuel temperature indicator is serviceable and there is power supply to the

indicator, the pointer should read between −40 to +57°C (−40°F to +135°F).

(7) Remove electrical power from the aircraft.

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LEGEND

1. Indicator.

2. Electrical connector.

3. Screw.

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FUEL

−40

TEMP

−40

am01_7333004_001.dg, kmw, aug14/2008

Fuel Temperature Dual Indicator − Removal and Installation

MASTER EFFECTIVITY:

Figure 201

73−33−00

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**ON A/C ALL

ENGINE INDICATING − GENERAL

1. General Engine indicating consists of torque (TRQ), high pressure compressor rotor speed (NH), low

pressure compressor rotor speed (NL), inter−turbine (ITT) temperature indicating systems for each engine. The indicators associated with these systems are located on the engine instrument panel in the flight compartment (refer to Figure 1). Sensors for the systems are shown in Figure 2.

Other indicators associated with engine operation and located on the engine instrument panel (Figure 1) are:

A. Propeller RPM (PROP RPM) indicating systems, refer to Chapter 61. B. Fuel flow (FF) and fuel temperature indicating systems, refer to Chapter 73. C. Oil pressure (ENG OIL PSI) and oil temperature (ENG OIL °C) indicating systems, refer

to Chapter 79. D. Fuel quantity (FUEL QTY) indicating, refer to Chapter 28. E. Fuel tank temperature (FUEL TEMP °C) indicating (Mod 8/0200 Aircraft refer to

Chapter 28)

All indicators incorporate white lighting, powered from the 5V dc lighting system. Warning lights incorporated for interturbine temperature (ITT) (located in interturbine temperature indicator) are interconnected with a dim and test control box. Refer to Chapter 33.

Torque (TRQ), interturbine temperature (ITT), fuel flow (FF), propeller RPM (PROP RPM), and high pressure compressor (NH) indicators are each provided with liquid crystal display (LCD) for digital readouts.

Pre mod 8−1/1101 aircraft indicators incorporate temperature controlled heaters which allow (LCD) operation down to −55 degrees C. in the flight compartment. Indicator heaters are powered from the 28V dc left and right secondary buses and circuit protection provided by (S2) and (A1) circuit breakers.

Mod 8−1/1101 aircraft indicators eliminate (LCD) heaters and heater power supplies therefore operation of the (LCD) readout is reduced to −15 degrees C. Analog operation of the servoed pointer is uneffected and will still function to −55 degrees C.

Two ENGINE FAIL indicator lights are powered to indicate engine failure on takeoff. They are located on the glareshield panels in the flight compartment, one in line with the pilot, and the other with the copilot. The lights are controlled by propeller autofeather system, and both lights come on if either engine fails. The autofeather system (refer to Chapter 61) is in operation only during takeoff.

Red flags on an engine condition panel are used to indicate faults or conditions leading to faults in fuel system filters, oil system filters, lube system chip detectors and electronic control units (ECU) (refer to Chapter 71).

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ENGINE INSTRUMENT PANEL

Figure 1

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ENGINE INDICATING − SENSING COMPONENTS ON ENGINE

MASTER EFFECTIVITY: See Pageblock 77−00−00 page 1

Figure 2

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PRE MOD 8−1/1101 LIQUID CRYSTAL DISPLAY (LCD) HEATERS − ELECTRICAL

MASTER EFFECTIVITY: See Pageblock 77−00−00 page 1

SCHEMATIC

Figure 3

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TORQUE INDICATING SYSTEM − DESCRIPTION AND OPERATION

1. General Refer to Figure 1. Two torque indicating systems are provided, one for each engine, to indicate the torsional

deflection (torque) between the reduction gearbox input shaft and the reduction gearbox layshaft.

Each system consists of a torque measuring system supplied with the engine, which is interconnected with an airframe supplied indicator and an engine supplied electronic control unit (ECU). The output of the measuring system is a dc voltage signal, which is directly proportional to sensed torque and a 5 volt reference voltage supplied by measuring system computing circuits. The signals are applied to the corresponding torque indicator, which provides torque readings in terms of percentage.

On Pre Mod 8/0886 aircraft, the position indicator ‘bug’ in each torque indicator is deactivated. Normal wiring is disconnected between ECU and ‘bug’. The position ‘bug’ is rewired to indicate zero torque.

On Mod 8/0886 aircraft, the position indicator ‘bug’ is not included with torque indicator.

2. Description and Operation Refer to Figure 1. A. Torque Measuring System

The torque measuring systems, each of which include a torque shaft and torque reference and displacement toothed wheels, a variable reluctance sensor (magnetic pickup) and a signal conditioner unit (SCU) are described in the Engine Maintenance Manual. For location of SCU and sensor, refer to ENGINE INDICATING − GENERAL, Figure 2. The SCUs of both engines are interconnected by electrical circuits to provide control in the propeller autofeather system (refer to Chapter 61).

B. Torque Indicators

Refer to Figure 2. The torque indicators located on the engine instrument panel, are powered from the left

and right dc essential circuit breaker panels and are protected by 5−ampere circuit breakers. ENG 1 TORQUE IND LEFT ESSENTIAL (K6) and ENG 2 TORQUE IND RIGHT ESSENTIAL (J5). Lighting for the indicators is white, powered from the 5V dc lighting system (refer to Chapter 33).

The indicators, each receive a 0 to 5V dc signal proportional to sensed torque from the corresponding SCU and provide a visual indication of engine torque, using a moving pointer against a fixed dial and an equivalent digital display. The indication covers a range of 0 to 120% torque.

On Pre Mod 8/0886 aircraft, the torque indicator ‘bug’ is inoperative and is driven to ‘zero’ torque setting by a negative bias produced from 26V ac left bus (refer to Figure 1, Sheet 1). On Mod 8/0886 aircraft torque bus is removed from torque indicator. The indicator dial is marked TRQ % and the scale on the indicator fixed dial is marked with major graduations in increments of 10% between 0 and 120%, with minor graduations in increments of 5%.

NOTE

: Torque values in excess of 120% can be read on digital display.

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Range marks are a green arc 0 to 90%, a yellow arc 90 to 105%, a dashed red limit (radial) at 105% torque and solid red limit mark (radial) at 90% torque.

On aircraft with SOO 8152 incorporated, range marks are a green arc 0 to 97.5%, a yellow arc 97.5 to 112.5%, a dashed red limit (radial) at 112.5% torque and solid red limit mark (radial) at 97.5% torque.

The digital display is a 4−digit liquid crystal display (LCD) type. Pre mod 8−1/1101 aircraft incorporate a heater associated with LCD operation, one in each indicator powered from the 28 V dc secondary buses, engine No. 1 indicator heater from the left bus and engine No. 2 indicator heater from the right bus (refer to ENGINE INDICATING

− GENERAL). A 0 to 5 V dc signal proportional to 0 to 100% torque is relayed to the flight data recorder

(refer to Chapter 31). An overtorque switch located in each indicator, is connected in series with overtorque

relay power supply circuits for the corresponding engine, relay 7712−K5 engine No. 1 and 7712−K4 engine No. 2.

In the event of engine overtorque, a circuit to ground is completed by the overtorque switch in the indicator. The overtorque and Np governing cancel relay energizes to connect the corresponding ECU ’cancel Np gov’ discrete to ’discrete return’ to cancel Np underspeed governing. The power uptrim relay of the opposite engine fuel control system energizes (in conjunction with autofeather circuits) to remove 28V dc power from the associated ECU ’opp. engine uptrim’ discrete, which initiates engine uptrim.

A press−to−test pushbutton on the front of the indicator can be used to verify operation of indicator internal circuits. Pressing the test button, with power applied to the indicating system causes the pointer to move to a position opposite a blue dot at 105% torque, with an equivalent reading on the digital display and initiates operation of the overtorque switch.

Overtorque switch operation is verified in conjunction with autofeather control circuits. With autofeather selected (refer to Chapter 61) a ground is established for the associated power uptrim relay for each engine. Pressing the press−to−test button on the front of the indicator results in dial indications, with overtorque switch actuation verified by a ‘power uptrim’ light on the engine instrument panel which turns on.

C. Torque Indicators − Functional Test

(1) Make sure 28V dc power on aircraft buses. (2) On left dc circuit breaker panel, make sure the following circuit breakers are closed:

ENG 1 TORQUE IND (K6) ENG 1 TORQUE COND (L6)

(3) On right dc circuit breaker panel, make sure the following circuit breakers are

closed: ENG 2 TORQUE IND (J5) ENG 2 TORQUE COND (K5)

(4) Make sure the power levers are at FLT IDLE positions. At the engine instrument

panel, push the AUTOFEATHER select switch/light (SELECT light on). Press and hold press−to−test switch on left indicator. Check that indicator pointer goes to blue dot at 105%, rim pointer (’bug’) goes to blue dot at 85% (Pre Mod 8/1292), digital display shows 105%.

MASTER EFFECTIVITY: See Pageblock 77−10−00 page 1

77−10−00

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De Havilland DASH8 100 Series Maintenance Manual

Specifications and Main Features

Frequently Asked Questions

User Manual