Q-AFM KEY NOTES

Question 0

Secondary Bus Failures

Answer 0

In the case of a secondary bus short, the overcurrent condition will immediately trip the associated TRU pri-mary circuit breaker. 7 s later, with the EPCU declaring TRU failed the contactor L to R Secondary bus tie will close, transferring the short circuit to the opposite side TRU. At that moment, the cross tie fuse is blown isolate ing the fault. This is indicated by a L TRU or R TRU caution light and loss of services on the associated sec-ondary bus.

Question 1

Main Bus Failures

Answer 1

If a main bus fault occurs, the EPCU prevents the upper horizonal and 2 vertical bus ties from closing, isolating the bus. The DC BUS caution light comes on to warn of the fault impending condition. If the fault persists after approximately 5 s, the EPCU sends a TRIP signal to the GCU, isolating the affected generator. The EPCU will also open and lock-out the contactors connecting the batteries to the affected main bus. The MAIN BATTERY or AUX and STBY BATTERY caution light(s) and related DC GEN caution light will come on as a result. The EPCU continues to monitor the operating buses. All main DC services on the faulted bus side will not function.

Question 2

Variable-Frequency AC Power

Answer 2

Two 115 V, 45 KVA AC generators (mounted on the propeller reduction gearbox) supply variable frequency (340 to 560 Hz) AC power. The AC power is supplied to the left and right AC buses. AC power sources are pre-vented from being operated in parallel.AC power is available once the condition levers are out of START & FEATHER in the MIN /850 to MAX / 1020 range and the GEN 1 and GEN 2 switches on the AC CONTROL panel are on.If one AC generator fails, the # 1 AC GEN or # 2 AC GEN caution light comes and the remaining generator is
capable of carrying the airplane’s AC electrical load except galley power. An automatic cross tie function, controlled by the AC GCU logic circuits, ensures that all variable-frequency buses are powered when only one AC generator is on line. Whenever a fault condition exists, the GCU of the inoperative generator issues a transfer request signal to the operational side AC GCU. The operational side AC GCU will issue a CLOSE command to the failed side line contactor. In this configuration, the remaining generator will power both AC buses. In this situation the load shedding relays will not allow power to the galley buses.The AC generators are protected from bus faults by the GCUs that detect any excessive load that might result from a short circuit on a bus. Once a heavy load is detected, the GCU isolates the bus and turns on the appropriate L AC BUS or R AC BUS caution lights.The # 1 AC GEN HOT or # 2 AC GEN HOT caution lights come on whenever an AC generator overheats. The AC generator must be switched off.

Question 3

Transformer Rectifier Units (TRU)

Answer 3

Located in the nose, the 2 Transformer Rectifier Units change 3 phases, 115 V, variable frequency AC input power into 28 VDC (300 A max) nominal output. The TRUs are unregulated but provide DC power in the range of 26 to 29 VDC during operation. Under normal conditions, each TRU powers its respective secondary bus. The L TRU or R TRU caution light comes on if either TRU is off line or failed. The L TRU HOT or R TRU HOT caution light comes on if the sensor in either detects an overheat condition. The light will go out when the temperature drops below the overheat condition. The 2 TRUs alone are capable of powering the entire DC system.

Question 4

Air Conditioning System

Answer 4

The Air Conditioning System (Figure 6.2-5) receives bleed air when the BLEED switches on the AIR CONDITIONING control panel (Figure 6.2-3) or the BL AIR switchlight on the APU CONTROL panel are selected on.The Air Conditioning System is controlled by selecting the CABIN and FLT COMP PACKS switches (Figure6.2-2) to the MAN or AUTO positions and then adjusting the temperature using the TEMP CONTROL knobs.These switch settings determine the bleed air source, manual or automatic Environmental Control System
(ECS) operation, and the air flow temperatures for the flight and passenger compartments. The ECS Electronic Control Unit (ECU) (Figure 6.2-6) controls the two Nacelle Shut-Off Valves (NSOV) to reg-
ulate the air flow to the air conditioning packs. The ECU receives bleed air pressure and temperature data from the pack inlet absolute pressure and inlet temperature sensors. The ECU uses these data to control bleed air flow through the pack Flow Control Shut-Off Valve (FCSOV). The ECU also uses this data to control bleed air
flow rate when APU bleed air is selected on

Question 5

Automatic Mode ( CPC )

Answer 5

When electrical power is first supplied to the system, a full power up self test is done. The FAULT alert light, on the Cabin Pressure Control (CPC) panel comes on momentarily during the power up test mode. If there is a failure in the system, the light will stay on. The system operation is fully automatic with the data programmed
into the controller (Figure 6.2-17).With the system in AUTO mode, a pre-programmed cabin pressure controller does all pressure scheduling from take-off to landing with minimal crew input. The computer receives inputs from the crew and various airplane systems, and modulates the aft outflow valve. This keeps a fixed schedule of cabin altitude versus airplane altitude for complete regulation of cabin pressure.

Question 6

On Ground ( CPC )

Answer 6

When the airplane is on the ground and the engine power lever angles are set at less than 60°, the aft outflow valve is positioned at the fully open position to prevent airplane pressurization. The aft safety valve located on the aft pressure bulkhead, also opens on the ground when at least one engine is running at idle, or the APU is operating.

Question 7

Take-off ( CPC )

Answer 7

When the engine power levers angles are set to greater than 60° the controller sends a signal to the aft outflow valve to modulate, as necessary, to provide two take-off sequences:
• Pre-pressurization
• Take-off abort
The aft outflow valve moves from the fully open position and starts to modulate to control the pressure changes that occur during take-off. After take-off (as sensed by the PSEU), the aft outflow valve modulates to keep the set airplane pressure.

Question 8

Pre-Pressurization ( CPC )

Answer 8

The purpose of automatic pre-pressurization is to avoid a cabin pressure “bump” at take-off. During this sequence the cabin is pressurized to 400 ft below the take-off altitude at a rate of 300 ft/min.
In the case of a take-off without bleed air selected, this sequence leads to both the aft outflow valve and the aft safety valve closing.

Question 9

Take-off Abort ( CPC )

Answer 9

The Cabin Pressure Controller (CPC) is in take-off mode for at most 10 minutes after lift off. This avoids the requirement to reselect the landing altitude in case of an aborted flight and emergency return to the departure airport. During 10 minutes after the take-off the pre-pressurization remains in effect as long as:
• The scheduled cabin altitude is higher than the theoretical cabin altitude, or
• The airplane altitude is less than the take-off altitude + 5000 ft (valid only for take-off altitude over 8000 ft)
Once one of the above conditions is met, the CPC begins flight scheduling.

Question 10

Fire Detection

Answer 10

When a fire overheat condition occurs, the alarm signals are processed by the Control Amplifier then sent to the Fire Protection Panel in the flight compartment. If a fire or overheat condition occurs in either engine, this will cause the gas within the APD to expand and turn on the following lights in the flight compartment:

• Applicable PULL FUEL / HYD OFF T-handle light (red) comes on
• Both ENGINE FIRE Warning Press to Reset lights (red) flash
• CHECK FIRE DET warning light (red) flashes
• Fire tone (optional)

Either ENGINE FIRE PRESS TO RESET indicator is pushed to turn the audible tone (optional) warning off and/or cancel the flashing engine fire lights. The ENGINE FIRE PRESS TO RESET stay on steady for the
duration of the alarm condition.

Question 11

Fire Extinguishing

Answer 11

The forward and aft bottle squibs are armed by pulling the PULL FUEL / HYD OFF handle. After arming, the extinguisher bottle is
discharged by selecting the EXTG switch on the fire protection panel to FWD or AFT position. An electrical signal is sent which ignites the Electro-Explosive Device (EED). When the EED explodes it ruptures a burst disc and the pressurized bottle then discharges the suppressant into the engine zones.

Question 12

Baggage Compartments - Smoke Detection and Fire Extinguishing

Answer 12

Fire extinguishing for the baggage compartments is performed by two High Rate (HR) fire extinguisher bottles and one Low Rate (LR) fire extinguisher bottle. Each baggage compartment has one high rate fire extinguisher bottle. The Low Rate fire extinguisher bottle is shared between the FWD and AFT baggage compartments but is located in the AFT equipment bay (rear fuselage).

Question 13

Fire Extinguishing-Baggage

Answer 13

Pushing the SMOKE / EXTG switchlight activates the High Rate fire extinguisher bottle into the aft baggage area. The AFT ARM light will go out and the AFT LOW light turn on. After a seven minute delay, the Low Rate fire extinguisher bottle will automatically discharge into the aft baggage area and the FWD LOW light will turn on when the LRD bottle has depleted.

Question 14

Engine Fuel Feed

Answer 14

Fuel to each engine is fed from the collector tank, from a primary ejector pump or an AC driven auxiliary pump and delivered to the engine driven pump (Figure 6.9-12). If the engine driven pump inlet pressure drops below a preset limit, the related # 1 or # 2 ENG FUEL PRESS caution light comes on. An AC (Variable Frequency) auxiliary pump in each collector bay serves as a back up source of fuel boost
pressure for take-off and landing and in case the related primary ejector pump does not supply the necessary fuel pressure. Related TANK 1 or TANK 2 AUX PUMP switchlights on the FUEL CONTROL TRANSFER panel control the auxiliary pumps manually (Figure 6.9-13).
A TANK 1 or TANK 2 AUX PUMP switch indicator on the MFD Fuel Page shows the position of the switchlight. When the pump is supplying sufficient boost pressure, the TANK 1 or TANK 2 AUX PUMP light on the Fuel Page will turn green and the related ON switchlight segment turns green. The engine feed shutoff valve closes when the related PULL FUEL / HYD OFF handle, on the Fire Protection Panel (FPP), is pulled (Figure 6.9-13). Advisory lights on the FPP show when the valve is open or closed. The fuel is filtered and heated by Fuel Oil Heat Exchanger (FOHE) before entering the FMU. If the fuel filter becomes blocked, fuel bypasses the filter. The # 1 or # 2 FUEL FLTR BYPASS caution light will comes on if a related bypass is impending.

Question 15

Fuel Transfer

Answer 15

Fuel can be transferred from one tank to the other to correct fuel imbalances or for fuel management. If the Fuel Quantity Computer (FQC) detects a fuel imbalance of more than 272 kg (600 lbs), a yellow [BALANCE] message flashes just above the FUEL legend of the ED. The message will flash until the imbalance is corrected. An imbalance condition will also be shown on the Fuel Page by the analog quantity dials changing to solid yellow. A TRANSFER switch on the FUEL CONTROL TRANSFER panel controls the fuel transfer system (Figure 6.9-14). When the TRANSFER switch is selected, the auxiliary pump in the donor tank operates automatically to pump fuel to the receiver tank. A signal from the operating pump causes the related ON switchlight segment to turn green. Electrically operated fuel transfer shutoff valves open for fuel transfer and close when the transfer
is stopped. Fuel transfer indications are also shown on the MFD Fuel Page. Once selected, fuel transfer will continue until deselected by the flight crew or until a high-level sensor in the wing tank which is receiving fuel detects an overfill condition, which automatically halts fuel transfer. The FUELING ON caution light is on if the refuel / defuel access door is open

Question 16

Hydraulic System Description

Answer 16

Main hydraulic power is provided by 3 independent hydraulic systems, designated # 1 (left), # 2 (right) and # 3 (aft) (Figure 6.10-1). The # 1 and # 2 hydraulic systems are normally pressurized by a single Engine Driven Pump (EDP) for each system. System pressure is maintained at 3000 psi. The # 3 hydraulic system is powered by an accumulator which is pressurized by a DC-Motor-Driven-Pump (DCMP). A pressure switch controls the DCMP operation to maintain the accumulator pressure within 2600 to 3250 psi. An electrically driven Standby Hydraulic Pump operates as a backup to the # 1 hydraulic system. It operates during the take-off and landing phases, or if # 1 engine fails.
A Power Transfer Unit (PTU) operates as a backup to the # 2 hydraulic system. The PTU is powered by the # 1 hydraulic system.If both engines fail, where both EDPs and the Standby Hydraulic Pump are unavailable, the DCMP in # 3 hydraulic system provides sufficient hydraulic power to the elevators for pitch control.

Question 17

Power Transfer Unit (PTU)

Answer 17

A Power Transfer Unit (PTU) operates as a backup hydraulic pressure to the # 2 hydraulic system. The PTU uses hydraulic pressure from the # 1 system to power a hydraulic motor (Figure 6.10-11). The motor
then operates a hydraulic pump to pressurize the # 2 system. Hydraulic fluid is not shared or transferred between # 1 and # 2 hydraulic systems during PTU operation. Hydraulic fluid must be available in the # 2 system for PTU operation.

Question 18

3 Hydraulic System

Answer 18

The # 3 hydraulic system is an independent system (Figure 6.10-12). The system operates automatically. During an emergency condition the left and right inboard elevator PCU’s are powered when the # 1 and / or # 2 hydraulic systems fail, or if a dual engine failure occurs.
The # 3 hydraulic system can also be engaged manually by pushing the HYD # 3 ISOL VLV switchlight on the HYDRAULIC CONTROL panel. Once pushed, an amber OPEN legend on the switchlight will turn on. An accumulator and an isolation valve are also installed in the # 3 hydraulic system. A 28 V DC Motor Driven Pump (DCMP) operates automatically to pressurize the accumulator and keep the accumulator pressurized between 2600 to 3250 psi. When the DCMP is not operating, the accumulator holds a reserve of pressure. The volume of the # 3 system reservoir is 2.6 qt (2.46 l).The DCMP operates intermittently and is controlled by 2 pressure switches installed on the accumulator isolation valve. One switch signals the DCMP to operate if system pressure drops to 2600 psi and commands the DCMP to turn off when system pressure reaches 3250 psi. The other switch turns on the # 3 STBY HYD PUMP caution light if system pressure falls to 900 psi, or the DCMP has been operating for longer than 60 seconds on
the ground. Electrical power is supplied to the DCMP by the standby battery.

Question 19

Accumulator Isolation Valve

Answer 19

The isolation valve is used in the # 3 hydraulic system to isolate the elevators from # 3 hydraulic system pressure. During normal flight operation, the system is in an active standby mode with the accumulator isolation valve (energized) closed. When open, the isolation valve allows hydraulic pressure from the # 3 hydraulic system to power the elevators (Figure 6.10-13). The isolation valve will open in flight if # 1 and / or # 2 hydraulic system pressure is lost, or, if # 1 and # 2 engines fail. The isolation valve can be manually opened when the HYD # 3 ISOL VLV switchlight is pushed, shown by an
amber OPEN legend on the switchlight. An additional pressure switch is installed downstream of the isolation valve. It turns on theELEVATOR PRESS caution light if # 1, # 2 and # 3 hydraulic systems are supplying pressure to all 6 elevator actuators. If the isolation valve malfunctions open, the # 3 hydraulic system will supply hydraulic power to the elevators, even though # 1 and # 2 hydraulic systems are operative. The ELEVATOR PRESS caution light will turn on. The OPEN legend in the switchlight will not turn on.

Question 20

Ice Detection System

Answer 20

There is no flight compartment control for the Ice Detection System (IDS). The system automatically operates as soon as 115 VAC power is available. The IDS uses 2 Ice Detector Probes (IDP) on the left and right side of the front fuselage (Figure 6.11-12). If either IDP detects more than 0.5 mm of clear ice, it is heated with power from the related 115 VAC bus. This de-ices the probe so that it can detect ice again.

If the REF SPEEDS switch is not selected to INCR and either IDP detects ice, an ICE DETECTED message will be flashing amber (yellow) in normal video on the ED just below the SAT indication.

If the REF SPEEDS switch is selected to INCR after either IDP detectes ice, the ICE DETECTED message willchange to steady white.

If the REF SPEEDS switch is selected to INCR before either IDP detects ice, then the ICE DETECTED message will be displayed in reverse white video for 5 s.

Selecting the REF SPEEDS switch to INCR will display an [INCR REF SPEED] message in white below the ICE DETECTED message confirming the Stall Protection System (SPS) has been modified for icing conditions.

The ICE DETECT FAIL caution light will come on if both ice detector probes fail. Failure of only one probe will
not cause the caution light to come on, as the system is redundant.

Question 21

Airframe De-icing System

Answer 21

Airframe de-icing can be controlled automatically or manually. Pneumatically actuated rubber de-icing boots are bonded to the leading edges of the wings, horizontal / vertical stabilizers and nacelle inlet lips (Figure 6.11-14). De-icing bleed air is taken from the bleed port of each engine and is available to inflate the boots regard less of the position of BLEED control switches. System pressure is regulated to 18 psi and shown on the DEICE PRESS indicator, located on the co-pilot’s side panel. An isolator valve interconnects the 2 systems. A BOOT AIR switch is used to control the isolator valve, which is normally open to ensure uninterrupted operation of either system if one engine is not operating. The ISO position can be used to check regulated pressure in each system individually or to isolate a system
leak. Regulated de-icer pressure is also used to inflate the forward passenger and aft baggage door seals and to operate ejector for the pressurization system AFT safety valve. The boots inflate and stay inflated, with pressurized air when the Dual Distributing Valves (DDV) are energized open. When not activated, boot ports are connected to suction to deflate and hold the boots flush with the leading edges. The AIRFRAME MODE SELECT rotary switch selects automatic de-icing, when set to SLOW (3 min) or FAST (1 min). The selector is self-homing such that a selection to SLOW or FAST and back to OFF will complete a full cycle. Automatic boot inflation sequence is controlled and monitored by the Timer and Monitor Unit (TMU) (Figures 6.11-13 & 6.11-14). The TMU controls the sequence and supplies a dwell period related to the selected rate (Table 6.11-1). Green WING, TAIL and nacelle inlet lip boot inflation lights show boot inflation sequence and confirm correct boot inlfation pressure.

Question 22

Engine Intake Heaters / Bypass Doors

Answer 22

An electric heater is installed in the intake flange of each engine. The heaters are powered by 115 VAC variable frequency and are energized when the engine intake bypass doors are opened. An oil pressure switch and temperature sensor in the heater control circuit prevents heater operation when the engine is shutdown and / or air temperature is above + 15°C. Heater operation is confirmed by the HTR segment on the ENGINE INTAKE switchlight coming on when the doors are opened.

Question 23

Landing Gear - Description

Answer 23

The main gear (MLG) retracts aft and has multiple disc brakes with an anti skid system (Figure 6.13-13). The nose gear (NLG) retracts forward and has steerable nosewheels (Figure 6.13-14). The landing gear (LG) is operated by the # 2 hydraulic system and is controlled by the landing gear selector lever on the LANDING GEAR control panel. There is an alternate (emergency) means of extension for the main and nose landing gear. Advisory lights give extension / retraction and fail / safe information. Each main gear has a pair of forward and aft doors hinged to the nacelle side structure (Figure 6.13-15). When the gear is up, all doors enclose the main wheels. With the main gear down, the forward door on each main gear stays open. The nose gear has a pair of forward and aft doors, which completely enclose the nose gear when the gear is up (Figure 6.13-16). With the gear down, the forward nose doors are closed, while the aft doors stay open. The Proximity Sensor Electronic Unit (PSEU) controls the landing gear, hydraulically operated gear doors and related advisory lights. It also monitors Weight-On-Wheels (WOW) sensors. WOW signals prevent gear retraction while on the ground. Failure of a WOW system turns on a WT ON WHEELS caution light. Redundancy is built in to ensure landing gear operation if there is a PSEU failure. An audible warning tone sounds, when the
gear is not down and locked with landing flap or power settings.
Ground lock pins are supplied for the main gear and an integral ground lock mechanism is controlled from outside the airplane for locking the nose gear. The main gear lock-pins may be kept in the forward compartment of the forward passenger door. With the gear extended, the pins are inserted into the main gear stabilizer brace assemblies (Figure 6.13-17). There are also landing gear door lock pins for the nose (Figure 6.13-18) and main (Figure 6.13-19) hydraulic doors. This prevent the hydraulic gear doors from closing

Question 24

Alternate Gear Extension

Answer 24

The landing gear extension INHIBIT switch is installed in the flight compartment ceiling, adjacent to the main LANDING GEAR ALTERNATE RELEASE door. Setting the switch to INHIBIT isolates all hydraulic pressure from the landing gear system. When the main LANDING GEAR ALTERNATE RELEASE door on the flight compartment ceiling is opened, it mechanically opens a bypass valve in the normal hydraulic extension system and gives access to the MAINL/G RELEASE handle. Pulling the handle releases the main landing gear doors and uplocks. The main gear will free fall but may not fully extend. The LANDING GEAR ALTERNATE EXTENSION door, on the flight compartment floor, must then be fully opened giving access to the alternate extension handpump and the NOSE L/G RELEASE handle. Opening the door mechanically operates the MLG alternate selector valve.If the MLG does not reach the down and locked position, the extension pump handle, located behind the co-
pilot, is inserted into the pump handle socket and operated to complete main gear extension and subsequent down lock (Figure 6.13-22). Both the LANDING GEAR ALTERNATE EXTENSION door and the MAIN LANDING GEAR ALTERNATE RELEASE door must be left fully open after alternate landing gear extension. When the NOSE L/G RELEASE handle is pulled, the nose gear uplock and doors are released and the nose gear free falls to a down and locked position, assisted by the airflow to a down and locked position. Illumination of the appropriate gear locked down advisory lights (green), in either the primary or alternate panel, is sufficient to conclude that the landing gear is down and locked.

Question 25

APU Fire Extinguishing

Answer 25

If a fire is detected, the APU automatically shuts down and the fire extinguishing agent is released after 7 seconds. If automatic fire extinguisher discharge fails, the BTL ARM light stays on. The guarded EXTG Switchlight can be pushed to discharge the fire extinguishing agent, if the BTL ARM is on.

Question 26

Flight Mode ( spoilers )

Answer 26

The spoilers operate in proportion to, the up going aileron to provide roll control. Turning either the pilot’s or copilot’s control column, operates the spoilers and ailerons at the same time. The # 1 hydraulic system powers the inboard spoilers and # 2 hydraulic system powers the outboard spoilers (Figure 6.8-21). At airspeeds greater than 170 KIAS, only the inboard spoilers operate, the Flight Control Electronic Control Unit (FCECU) disables the outboard spoilers. At decreasing airspeeds less than 165 KIAS, inboard and outboard spoilers
operate. If the outboard spoilers are not disabled above 185 KIAS or activated below 150 KIAS, the SPLR OUTBD caution light turns on.Pushing either SPLR1 or SPLR2 switchlight, inhibits hydraulic pressure to its related spoiler PCU extend ports. This turns on the ROLL SPLR INBD HYD or ROLL SPLR OUTBD HYD caution light. The continuous hold down pressure returns the related spoilers to the down position.

Question 27

Ground Mode ( Spoilers )

Answer 27

There are 2 lift-dump valves in the inboard spoiler system and 2 in the outboard spoiler system for ground spoiler operations. The lift-dump valves in each spoiler system, are hydraulically in series; both valves mustopen together before the spoilers can extend on the ground. When the lift-dump valves are energized open, hydraulic input commands are sent to the PCUs which fully extend both inboard and outboard spoilers.The lift-dump valves are energized by signals from the FCECU and the Proximity Sensor Electronic Unit (PSEU). For the spoilers to extend on landing, the FCECU and PSEU must receive valid input signals before energizing the lift-dump valves (Figure6.8-22).

Question 28

POWERPLANT

Answer 28

The Dash 8-Q400 is powered by two Pratt & Whitney PW150A turboprop engines. Each engine drives a six bladed, constant speed, variable pitch, fully feathering Dowty R408 propeller through the engine gearbox. The powerplant develops 4,580 Shaft Horse Power (SHP) under normal take-off conditions. An automatic uptrim on a manual MTOP rating selection allows either engine, to develop a maximum take-off power of 5071 SHP, for a brief period of time, if an engine failure occurs during take-off.

Question 29

Accessory Gear Box

Answer 29

An accessory gearbox mounted on top of the engine is driven by the high pressure compressor rotor NH, and operates:
• Oil Pressure and Oil Scavenge Pumps
• High Pressure Fuel Pump
• Permanent Magnet Alternator (PMA)
• DC Starter / Generator

Question 30

Permanent Magnet Alternator

Answer 30

The primary source of electrical power for the engine control system is the engine mounted Permanent Magnet Alternator (PMA). The PMA has independent coils that provide electrical power to the individual channels of the FADEC when gas generator speed (NH) is above 20% minimum. The aeroplane essential power busses provide alternate electrical power to the FADEC for engine starting and in the event of a PMA malfunction.

Question 31

Automatic Take-off Power Control System (ATPCS)

Answer 31

During an engine take-off, an Automatic take-off Power Control System (ATPCS) augments the power of the engine, without pilot intervention, in response to a loss of power of the opposite engine. This function is also referred to as “Uptrim”. The working engine’s FADEC will respond to the Uptrim signal from the failed engines
PEC/AF unit by changing engine rating from NTOP to MTOP.
The working ATPCS is armed when both PLAs are “high” and local torque engine is “high”. If an engine fails (i.e. engine torque is “low”) an Uptrim signal is commanded by the failed engine PEC to theworking engine FADEC. The working engines power is increased 10%.
An Uptrim condition is indicated to the pilot by:
• the UPTRIM indication on the ED
• a change in the engine rating from NTOP to MTOP
• a change in the torque bug from NTOP to MTOP

Question 32

Engine Emergency Power Control

Answer 32

Engine Emergency Power is achieved by advancing the Power Levers above the Rated Power detent position, in the overtravel region. Both the engine control and propeller control systems are affected. In the over travel region, engine power increases linearly from Rated Power to 125% of MTOP, propeller speed increases to 1020 RPM regardless of Condition Lever setting. This speed is latched until Power Levers are returned to the Rated Power detent, or below, and the Condition Levers are selected to MAX/1020, or START/FEATHER or FUEL OFF.

Question 33

Torque Limiting

Answer 33

The Torque Limiting Logic in the FADEC prevents engine torque from exceeding a given threshold which is function of PLA and ambient conditions. Generally torque is limited to 35% in reverse, 106% in the forward power range, and 125% in the overtravel range. However, during such events as caused by a spurious feathering of the propeller at high power, the transient overtorque can exceed this steady state threshold. The FADEC uses anticipation in this control loop to rapidly
reduce NH to prevent overtorque in exceedance of 135%

Question 34

ENGINE START SYSTEM

Answer 34

Engine starting is accomplished using the STARTER/GENERATOR in conjunction with the ignition and fuel control systems. The STARTER/GENERATOR rotates the High Pressure (NH) compressor through the accessory gearbox, to develop the necessary airflow and engine RPM before fuel is introduced. The start system is armed by selecting the engine to be started on the ENGINE START panel and turned on using the switch. The start sequence is initiated by pressing the START switchlight and selecting the condition lever to the START &FEATHER position at the first indication of NH. Once the start has been initiated, the FADEC controls the starting sequence in the following manner:

• When the STARTER/GENERATOR has increased NH speed to 8%, the FADEC commands ignition on and schedules fuel flow as a function of NH, ambient temperature and ambient pressure.
• Only one of the two ignitors is turned on (this is to identify any failures in the dual channel ignition system). If the engine does not light-off within 8 seconds of fuel flow being selected on, the FADEC turns on both ignitors, and starts a count towards logging a fault against the faulty igninition. Light-off is defined as anincrease of 20°C in ITT.
• When NH is greater than 50%, the ignitor(s) is automatically turned off.
• The FADEC controls engine run-up to the requested NH speed or power. During ground starts, to ensure that the engine start does not cause overtemperature, the FADEC has active ITT limiting to reduce the fuel flow if required (below the standard start schedule).

Question 35

Rudder Jam

Answer 35

If a jam occurs in a rudder PCU, the corresponding RUD 1 or RUD 2 PUSH OFF switchlight turns on. The illuminated RUD 1 or RUD 2 switchlight must then be pushed to depressurize the affected PCU. The PUSH legend will go out and the OFF legend will remain on as a reminder that the switchlight has been pushed OFF. The # 1 RUD HYD or # 2 RUD HYD caution light will turn on as the PCU is depressurized. The FCECU will reschedule the regulated hydraulic pressure to the operative PCU to maintain the required rudder authority. As directed by paragraph “Rudder Actuator Malfunction” of the AFM - “Abnormal Procedures”, only one RUD PUSH OFF switchlight shall be pushed at a time. If both RUD 1 and RUD 2 PUSH OFF switchlights are pressed inadvertently, the OFF legend will go out, both RUD 1 and RUD 2 PUSH legends will turn on and the
previously depressurized PCU will be re-pressurized. This ensures the rudder control system remains powered. Pushing the non-jammed switchlight again turns out both PUSH legends, de-pressurizes the jammed PCU and turns on the appropriate OFF legend on the jammed side. If instead the jammed side switch is pushed, the jammed side RUD PUSH OFF light will turn on while the non-jammed side will be depressurized and its corresponding OFF legend will turn on.
When the aircraft is parked on the ground with engines not running, one or both RUD 1 and RUD 2 PUSH OFF switch lights may be illuminated under conditions of strong tailwinds. This is a result of the rudder PCU bungees being compressed when the rudder is moved to one side under the influence of the wind. As soon as hydraulic
pressure is available to the PCU following engine start, the rudder will center and the RUD PUSH OFF switchlights will go out.

Question 36

Starting from DC External Power

Answer 36

A Ground Power Unit (GPU) can be connected to the DC external power receptacle (Figure 6.5-18) on the left side of the forward fuselage. The GPU can supply DC power for the airplane EPGDS and for engine starting (Figure 6.5-17). Contractors connecting the main and aux bateries to the main buses are closed automatically
upon selection of engine start, as the main and auxiliary batteries assist in the start. The standby battery is diode isolated from the left main, to ensure an acceptable level of voltage on the ESS Busses during the starting process. When the starting process is terminated the power source is still the external power. While DC external power
is connected to the airplane, the generators connections to the main buses is inhibited by the EPCU. The main and auxiliary batteries stay connected to the main buses and the standby battery is reconnected.
Once the DC EXT PWR switch is turned off after engine start, the generators will come on line if the GEN switches are in the ON position. Both vertical bus ties connecting the main to the secondary buses will remain closed until DC power is available from the TRUs.
The EPCU incorporates external DC power protection from too high or too low supply of external DC power voltage. If the external voltage is more than 31+ 0.5 / - 0.75 VDC, or less than 22 ± 1 VDC, an over / under voltage condition will cause the external ground power to stop supplying electrical power to the airplane. If the external power over / under voltage is rectified, the external power source can be reselected by moving the DC EXT PWR switch to OFF and then to EXT PWR.

Question 37

Pitch Control System

Answer 37

Pitch control of the airplane is maintained by 2 mechanically controlled and hydraulically powered elevators (Figure 6.8-23). The elevators are attached to the trailing edge of the left and right horizontal stabilizers. The left control column operates the left elevator and the right control column operates the right elevator. However both control columns are connected to each other by the pitch disconnect system so that they both operate together. Fore and aft movement of the pilot’s and co-pilot’s control columns is transferred through 2 fully independent cable and pulley control circuits to the elevator Power Control Units (PCU).
There are 3 identical hydraulic PCUs (outboard, centre and inboard) on each elevator. The outboard and centre PCUs on each elevator are active at all times while the inboard PCU is a standby. The # 1hydraulic system supplies power to the left and right outboard PCUs. The # 2 hydraulic system supplies power to the left and right centre PCUs. The standby # 3 hydraulic system supplies power to the left and right inboard standby PCUs when required. The HYD # 3 ISOL VLV pushbutton on the HYDRAULIC CONTROL panel when pushed, manually activates the inboard PCUs. This will cause the ELEVATOR PRESS caution light to turn on if the # 1 and # 2 hydraulic systems are functioning. The # 3 isolation valve will also activate automatically when # 1 and / or # 2 hydraulic system fails. Pitch trim is accomplished by two pitch trim actuators. The actuators are controlled automatically by the autopilot or manually by the trim switches on the pilot’s and co-pilot’s control column. Elevator trim position is shown
on the elevator trim indicator located on the left side of the centre console. If a mismatch occurs between the left and right elevator an ELEVATOR ASYMMETRY caution light comes on. Elevator position indication is displayed on pilot’s Multi-Function Display (MFD). Gust protection for the elevators is supplied by trapped hydrau-
lic fluid within the actuators when the system is depressurized.