This is a hyperlinked index to the text only portion of the significant text of the FSI manual corrected slightly where necessary
The Environmental Control System (ECS) consists of the air conditioning pack and its Electronic Control Unit (ECU).and uses bleed air from the engines or Auxiliary Power Unit (APU) to supply conditioned air to the flight compartment and cabin.
This pack consists of two Air Cycle Machines (ACM), that are integrated with single primary and secondary heat exchangers. They are located in the aft fuselage equipment bay. This configuration provides the redundancy of two ACMs supplying both cabin and flight compartment.
The air conditioning system receives bleed air when the BLEED switches on the AIR CONDITIONING control panel, or the BL AIR switchlight on the APU CONTROL panel, are selected on. Selecting the CABIN and FLT COMP PACKS switches to the MAN or AUTO positions, and then adjusting the temperature using the TEMP CONTROL knobs control the air-conditioning system. These switch settings determine the bleed source, manual or automatic Environmental Control System (ECS) operation, and the airflow temperature for the flight deck and cabin.
The OFF selection for both packs closes the pack inlet FCSOV and TURBINE SOVS , bleed air does not flow.
When selecting only one pack to MAN or AUTO, the ECS controller opens
the pack inlet FCSOV,
the appropriate Turbine SOV and
both pack Bypass Valves.
The recirculation fan will run at low speed.
When selecting both packs to MAN or AUTO, the system will operate normally.
The recirculation fan will run at high speed.
If a digital channel becomes inoperative, the associated Turbine SOV will revert to the
associated analog control. The remaining digital channel will operate the pack FCSOV to adjust the required flow.
Both digital channels of the ECU share control of the pack FCSOV. During flight, one digital channel gets full control of the pack FCSOV (the other channel gets full control during the next flight).
The digital channel in control modulates the nacelle SOV to set flows to the packs, and the flight deck and cabin based upon:
-Performance required by temperature selections
-Mass bleed flow measured at wing duct
-Air source (single bleed, dual bleed or APU)
The analog backup channels do not modulate the pack FCSOV. They control the valve to either fully open or close.
If a malfunction occurs to the pack FCSOV, it defaults pneumatically to the open position to permit continued ECS operation. However, if both digital channels of the ECU lose electrical power or fail, the pack FCSOV defaults to the closed position. ECS operation stops and the packs shut off. If this occurs, air must be supplied to the cabin and flight compartments using emergency ram air ventilation.
If one engine is shut down, the operating digital channel regulates bleed airflow with the pack FCSOV.
With both packs operating, the left digital channel uses approximately half of the air from the left pack for the flight compartment temperature. The right digital channel uses the other half of the air from the left pack and all of the air from the right pack to control the cabin compartment temperature. Therefore, the cabin receives approximately 75% of the total airflow.
The recirculation fan draws cabin air through the recirculation filter, mounted behind the aft baggage compartment. The air is then mixed with pack conditioned air. The RECIRC fan switch, on the AIR CONDITIONING control panel, controls the on/off operation of the recirculation fan, independent of the ECU. However, the ECU will use the operating conditions and the length of time the fan was on to determine the fan speed.
When the switch is set to the RECIRC position, the fan starts at low speed (to reduce initial current draw), then changes to high speed. The fan operates at low speed if one pack is off.
The temperature control and indication system is controlled from the AIR CONDITIONING control panel. The Electronic Control Unit (ECU) is the interface between the AIR CONDITIONING control panel and the mechanical and electrical components of the air conditioning system.
The ECU has dedicated cabin and flight compartment duct temperature sensors in order to regulate the supply between 2.8° C and 71° C The minimum temperature of 2.8° C makes sure that there is no ice formation on the condenser.
Each duct supply has overtemperature sensors.
The ECU also has cabin and flight compartment zone sensors to maintain the cabin and flight compartment temperatures between 15° C and 27° C.
Separate Flight and Cabin duct and Cabin zone sensors supply information to the cockpit gauge on the AIR CONDITIONING panel.
When the PACKS switches are set to AUTO, the digital channel in control opens the pack FCSOV and the turbine shutoff valves. The ECU modulates the bypass valves to add warm air to the cool air from the ACM’s. The ECU controls the pack outlet temperature based on the settings of the CABIN and FLT COMP temperature knobs, and information from appropriate zone temperature sensor in order to provide the desired temperature.
When the PACKS switch is set to MAN, or if the digital channel in control fails, the analog channel supplies temperature control. Control is identical except that temperature control is based on feedback information from the supply duct temperature sensors.
When both PACKS switches are set to OFF, this closes the pack FCSOV and the turbine shut-off valves, shutting down the packs.
The FLT COMP temperature knob controls the flight compartment temperature. A flow control lever is located under the left and right side windows on the sidewall. The levers regulate the quantity of air flowing to the flight compartment. Pilots should understand that without airflow there is not any temperature control.
The CABIN temperature knob has a switch at the full counter clockwise F/A position. Turning the knob to the F/A position turns on the FA CONTROL ENABLED light on the attendant's control panel. Both the FA CONTROL ENABLED light, and the cabin temperature gauge, on the attendant's control panel operate independently from the ECU.
Air supply to the flight compartment is ducted from the air conditioning pack, through the rear pressure bulkhead. The duct is then divided so that the left side supplies flight compartment air, while the right side supplies cabin air. . The air is then ducted under the baggage compartment floor forward under the cabin.
Before reaching the flight compartment, the distribution system also supplies conditioned air to the aft baggage compartment inlet (???), forward lavatory gasper and the forward flight attendant's gasper.
Each duct has a temperature sensor, and an overtemperature switch
At the flight compartment bulkhead, the duct splits, one for the left side and the other for the right side. Each side has upper level demist nozzles for the pilot's and copilot's side windows. The lower level outlets include a foot warming piccolo tube (near the rudder pedals), a fixed grille near knee height and a large torso gasper. A small manually controlled gasper at window height is also provided.
Two flow control levers located at shoulder height regulates the quantity of air flowing to these various outlets.
The air is ducted under the baggage compartment floor, where it splits into an upper and lower supply duct for each side of the cabin The upper distribution duct supplies the Passenger Service Unit (PSU) gaspers and the sidewall downwash and ceiling upwash vents. The lower distribution duct supplies the dado panels.
A distribution damper controlled by the ECU using the signal from the cabin zone temperature sensor directs air to the upper or lower duct, as conditions require.
I.e. During heating operations, most of the warm air goes to the lower cabin dado panels and during cooling operations, most of the cool air is directed to the PSU panels and gaspers,
The right digital channel of the ECU controls the electric motor of the distribution damper. If the right digital channel or the electric motor fails, the damper valve will remain in its last position.
The aft baggage compartment has an inlet and outlet ventilation valve. They are normally open but close when smoke is detected in the baggage compartment and/or when electrical power is lost. Two white advisory lights on the Fire Protection Panel come on when the inlet valves are closed.
The digital channel in control will shut off the pack flow control and shutoff valve (and stop pack operation) if:
• Both PACKS switches are set to OFF.
• The Built In Test (BIT) function of the ECU detects an unacceptable condition.
• Two or more of the three pack overtemperature protection switches detect an overtemperature condition.
If a pack overheat condition occurs, the
Related pack shuts down
Related CABIN PACK HOT or FLT COMPT PACK HOT CAUTION light comes on.
The circuit latches and the light will remain on.
If a flight compartment or cabin duct supply overheat condition occurs, the related:
• FLT COMPT DUCT HOT caution light comes on
• CABIN DUCT HOT caution light comes on.
The ECU will modulate the related pack bypass valve toward closed to aid in cooling the duct. The related FLT COMPT DUCT HOT or CABIN DUCT HOT caution light latches, and will stay on until the related PACKS switch is set to OFF (??)
If both packs are shut down cabin airflow and pressurization will be lost. During unpressurized flight the cabin and flight compartment can be ventilated with outside ram air. (See Chapter 18 Pressurization).
The avionics cooling system has three fans. The system removes hot air from the avionics equipment, and the five Liquid Crystal Displays (LCD).
Control of the system is completely automatic. There are three identical fans. Only two of the three fans are required to be operational for dispatch.
The hot air is exhausted under the floor behind the flight compartment. Each duct assembly alone can supply enough cooling for continuous operation of the LCDs. 2 Fans run whenever the electrical power is applied to the aircraft DC main bus. If either of these fans fail, the standby fan (Fan 3) automatically starts operating.
When only the battery power is available, 1 Fan only operates in low speed. The standby fan is not available when operating on battery power. This operational mode is capable of supplying the minimum airflow required for the LCDs to operate at reduced brightness.
A fan operating at Low Speed Mode (LSM) can supply enough airflow to meet the avionics and LCD cooling requirements while still meeting the battery loading
Failure of any fan is recorded in the Central Diagnostic System (CDS). There is no indication of a single fan failure to the flight crew. If two fans fail, a FANS FAIL message is shown in white on the Engine Display (ED). If the aircraft is on the ground. Maintenance is required before flight.
The ducting for each avionics cooling fan has zone temperature switches located under the LCDs. These switches inhibit the fan operation on the ground, when the flight compartment temperature is below 5° C (Fans Fail Message). The temperature switches are disabled when the aircraft is airborne.
The optional Auxiliary Power Unit (APU) consists of a gas turbine engine driving a DC starter-generator. It supplies bleed air to the Environmental Control System (ECS), and 28 VDC to the electrical system. The APU operates on ground only. Two doors on the bottom of the tail cone access the APU. Aircraft with the APU have a titanium tailcone and firewall.
The start control, normal operation, and malfunction monitoring is automatically performed by the APU FADEC. The APU start can be powered from either the aircraft batteries or external power. Intake air is drawn through a screened inlet duct on the right rear of the fuselage. Exhaust gases are discharged through an upward pointing outlet at the aft end of the titanium tailcone. The APU is protected by its own automatic fire detection and extinguishing system that continuously monitors the APU and its compartment whenever electrical power is supplied to the system. The APU control panel is mounted on the overhead console, in the flight compartment.
APU fuel is supplied from the left wing collector bay through an APU shutoffvalve. A gravity-fed, APU-driven fuel pump keeps positive fuel pressure to the APU engine.
The APU shutoff valve opens when the APU PWR switchlight is pushed and closes when the APU is shutdown. Fuel is automatically scheduled for starting, acceleration, and speed regulation.
The APU shutoffvalve will close if:
• PWR switchlight is pushed off
• FIRE TEST pushbutton is pushed
• APU automatically (FADEC) shuts down caused by;
Fire bottle depleted
Aircraft airborne (main wow nose on earlier ac))
The APU air intake inlet is on the top right side of the tail cone. There is a wire mesh grill to keep foreign objects out, and an eaves trough for possible fluid leaks.
An optional APU inlet louvered cover can be installed help keep out snow and contaminates. It must be removed to operate the APU when temperatures are above 25 Deg. C.
The APU compartment is ventilated by the APU exhaust air ejector system.
The APU has a 28 VDC starter‑generator). An start requires either aircraft batteries or external power. When starting, the starter stays engaged until the APU reaches half its operating speed. When the APU is operating, the RUN segment comes on to show that the generator mode is available to supply 28 VDC. The WARN segment of the GEN switchlight comes on when the generator is not supplying power.
The APU electrical load and voltage can be monitored on the ELECTRICAL page. ‑
With main engine DC starter-generators are on line, the APU generator will continue to supply power in parallel to the DC buses. If DC external power is applied to the buses, the APU generator will go off line. (WARN amber)
The starter‑generator is air-cooled.
If a starter‑generator fault is detected, the:
-Starter‑generator is disconnected from the right main feeder bus
-ON segment of the GEN switchlight goes off.
-WARN segment of the GEN switchlight goes amber
-APU caution light comes on.
If the starter‑generator overheats, in addition to the above,
- GEN OHT advisory light comes on.
When the APU is operating, pushing the BL AIR switchlight on the APU control panel opens the APU bleed air valve, the OPEN segment (green) of the BL AIR switchlight illuminates, bleed air is supplied to the aircraft, and the Cabin Pressure Control System (CPCS) aft safety valve is held open.
The APU FADEC controls the APU bleed air valve. If the APU exhaust temperature is above limits, the APU FADEC reduces the bleed air supply. This gives APU generator load priority over bleed air.
Two check valves prevent APU bleed air from entering the engine bleed air supply, including airframe deicing.
If either aircraft BLEED air toggle switch is set to 1 or 2:
• APU bleed air will not open
If an APU bleed problem is detected: i.e. FLT COMPT DUCT HOT, CABIN DUCT HOT, FLT COMPT PACK HOT, or CABIN PACK HOT caution light(s)
• APU bleed air will not be supplied (off)
APU fire protection is by a control amplifier and an Advance Pneumatic Detector (APD). The system monitors the APU hot section and exhaust whenever the right essential 28 VDC bus is powered. It is operated from the Fire Protection Panel (FPP) on the overhead console. The system has a stainless steel fire extinguisher bottle and distribution tubing, and an APD routed along the tailcone above the APU.
When a fire or overheat condition is sensed, the:
• FIRE light (red), on the FPP, comes on
• CHECK FIRE DET warning light flashes (red)
• MASTER WARNING light flashes (red)
• BTL ARM advisory light (amber); then (out) if the bottle has discharged
• EXTG switchlight (white)
• APU FUEL VALVE CLOSED advisory light (white)
• APU FUEL VALVE OPEN advisory light (out)
• FAIL light (amber)
• APU caution light (amber)
• MASTER CAUTION light flashes (amber)
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 guarded EXTG switchlight can be pushed to discharge the fire-extinguishing agent, if the BTL ARM is on.
NOTE once the bottle has been discharged, restarting the APU is prevented until the bottle has been recharged.
If the APD signal, to the control amplifier, opens because of a failure or power loss, the FAULT segment will come on, and the CHECK FIRE DET warning light will flash.
If the control amplifier APU circuit fails, then only the FAULT segment will come on.
If an APU FADEC detects a fault the:
• APU automatically shuts down
• FAIL segment of the APU PWR switchlight (amber)
• APU FUEL VALVE CLOSED advisory light (white)
• APU FUEL VALVE OPEN advisory light (out)
• APU caution light (amber)
The APU PWR switchlight must be reselected after an automatic shutdown or failure to start.
NOTE Do not restart the APU after an automatic shutdown, if the FIRE advisory light is on.
Push the GEN off, Push bleed off. Push the PWR switchlight.
Note: A failed start forces a 70-sec wait as per limitations.
The Electrical System provides for energy conversion, distribution, storage, control, protection, monitoring, and indication and has Direct Current (DC) and Alternating Current (AC) power systems. The DC power system includes a battery system. There are external connections for DC and AC external power.
Three NiCad batteries, two engine driven starter/generators, two Transformer Rectifier Units (TRUs) and an Auxiliary Power Unit (APU) supply the DC power system. The TRUs supply 28 VDC and are powered by two engine driven 115 VAC generators.
Electrical power is distributed by a bus system that reconfigures for power and bus failures, automatically.
All AC and DC aircraft services can be powered from the AC generators or the AC external power alone.
The Electrical Power Generation and Distribution System (EPGDS) has an Electrical Power Control Unit (EPCU) to control, monitor and distribute DC power to the aircraft's electrical buses. The EPCU automatically reconfigures the EPGDS for power source and bus failures, by the closing and opening of bus ties contactors. Contactor control is determined by automatic functions during the operation of the aircraft. Requests are achieved through the selection of switches in the flight compartment that may be vetoed by the EPCU.
System parameters are monitored on the Engine and Systems Integrated Displays, in the flight compartment. Turning the related Multi Function Display (MFD) selector, on the Engine and Systems Integrated Displays Control Panel (ESCP), to SYS (System), shows electrical system data on the ELECTRICAL page.
The DC generation system has the following sources:
- Main, Auxiliary, and Standby Batteries
- Two Starter/Generators
- Two Transformer Rectifiers Units (TRU)
- DC External
- APU Starter/Generator
The power sources supply power to the following buses in order of priority:
- Left and Right Essential
- Left and Right Main
- Left and Right Secondary
The main and auxiliary 24‑volt NiCad batteries are located in the lower left nose compartment. The main and auxiliary batteries have a 40amp hour capacity. The standby 17amp/hour battery (optional 40 amp hr) is located in the upper left nose compartment, The MAIN and AUX batteries are used for engine cranking. (assist in the case of GPU) In flight all three batteries ensure backup power to the aircraft essential services for 45 minutes.
Setting the BATTERY MASTER switch to the BATTERY MASTER position connects all three batteries to the essential buses. The connection is hard wired, and independent of the EPCU operation. The EPCU itself is energized from the essential buses.
NOTE; Battery power cannot be applied to the secondary buses.
Each battery has a dedicated switch. MAIN BATT connects the main battery to the right main bus. AUX BATT connects the auxiliary battery to the left main bus.
The connections of the main and auxiliary batteries to the main buses are controlled by the EPCU logic.
-Depending on switch position
-Emergency operation (defined later)
-Bus fault detected (defined later)
STBY BATT connects the standby battery-charging relay to the left main bus. This connection is done electronically by the EPCU if:
- DC external power connected
- No emergency operation
- No bus faults detected
MAIN BATTERY, AUX BATTERY and STBY BATTERY caution lights come on when the related battery is not connected to its feeder bus.
The Main, Aux and Stby battery temperatures are shown on the MFD ELECTRICAL page. If the battery temperature goes above normal limits, the related temperature digital value will be shown in yellow. If the temperature increases further, the digital values are shown in red. If the battery temperatures get too hot, the related MAIN BAT HOT, AUX BAT HOT or STBY BAT HOT warning light comes on. The related warning light will go off when the temperature drops below the overheat condition.
The starter/generators are located on the accessory gearbox of each engine. Each starter/generator serves as a starter motor, until the engine speed is at 50% NH. It then changes to generator operation, if the DC GEN switch is in the 1 or 2 position, following a successful engine start. Each generator output is monitored and controlled by its Generator Control Unit (GCU). After engine start, the GCU makes sure the generator supplies 28.5 VDC (400 Amps max) to its feeder bus regardless of load. The EPCU will then monitor and control the system.
If the DC generator is disconnected from its feeder bus, the #1 DC GEN or #2 DC GEN caution light to come on.
If a generator overheats, the related #1 DC GEN HOT or #2 DC GEN HOT caution light will come on. This light will go out when the generator temperature drops below limits.
Two TRUs, located in the nose, change three-phase, 115 V variable frequency AC input power into 28 VDC (300 Amps max) nominal output. The TRUs provide DC power in the range of 26 to 29 VDC.
Under normal conditions, each TRU powers its related secondary bus. If either TRU is off line, the related L TRU or R TRU caution light comes on.
In an overheat condition, the L TRU HOT or R TRU HOT caution light comes on. The light will go out when the temperature drops below limits.
The two TRUs alone can power the entire DC system.
The APU, controlled by the APU GCU, supplies 28 VDC to all the DC buses. For APU starting the APU Starter/generator is connected to the right main feeder bus. After the APU is started, the starter/generator is available to supply power in parallel with the main and auxiliary batteries to start the aircraft engines.
The APU shuts down when main gear is not weight on wheels (WOW).
The DC Power subsystem is tolerant to power source failures. It has the two generators and two
TRUs sources of DC power in flight. When all DC power sources are operational, each source powers its own dedicated bus.
If a failure is detected, the EPCU will cause the bus system to reconfigure the power flow, to make sure the flight is completed with minimal effect.
Four bus tie contactors, controlled by the EPCU, connect the appropriate feeder buses together, when there are one or more DC power sources not operating. Automatic load shedding will occur only if more than two DC sources fail.
For example, if one DC generator fails, the L to R Main bus tie closes and the other DC generator then powers both main feeder buses.
If one TRU fails, the L to R Secondary bus tie closes, and the other TRU then powers both secondary buses. If both DC generators fail, the vertical bus ties close, and the TRUs supply DC power to both the secondary and main buses. If both TRUs fail, the vertical bus ties close, and the DC generators supply power to both secondary buses.
The EPCU operates in emergency mode if:
-the airplane is airborne
-both DC Generators are failed and
-at least one TRU failed
And automatically disconnects the batteries from the main buses.
NOTE The main buses are not powered in this mode
If an engine start attempt is made during emergency operation mode, the batteries will automatically be reconnected to the main buses for the duration of the start attempt, regardless of BATT switch selection.
The EPCU and the DC GCUs circuitry protects the DC system against faults (short circuits and overloads) on the main and secondary buses.
If a main bus fault occurs (GCU senses excessive gen load), the EPCU isolates the bus es (no bus ties). The DC BUS caution light comes on to warn of the fault. If the fault persists for 5 seconds, the GCU isolates the affected generator. The batteries are isolated from the affected main bus. The MAIN or AUX BATTERY and STBY BATTERY caution lights and related DC GEN caution light will come on. All main DC services on the faulted bus side will not function.
Manual operation of the main bus tie through the MAIN BUS TIE switch is not possible once the EPCU has reacted to a fault.
If the fault subsequently clears, or faulty generator/battery source is isolated, power may be restored with the BUS FAULT RESET switch.
If a secondary bus short occurs, the overcurrent condition trips the related TRU circuit breaker. The EPCU closes the L to R Secondary bus tie contactor, transferring the short circuit to the opposite side TRU. The crosstie fuse is blown (it has a lower value than the TRU circuit breaker), isolating the fault.
A L TRU or R TRU caution light, and loss of services on the related secondary bus indicate this situation.
Starting is initiated by selecting an engine position using the SELECT switch on the ENGINE START control panel. The starter is engaged by pushing the START switchlight. The starter/generators are powered from the main buses only. The power source can be;
DC external power
In all cases the BATTERY MASTER, MAIN BATT, AUX BATT, STBY BATT and MAIN BUS TIE switches are selected on before start.
During a battery start, only the main and auxiliary batteries, in parallel, participate in the starting process. The standby battery is isolated from the left main bus by a diode to ensure there is an acceptable level of voltage on the essential buses.
During start the EPCU opens the contacts between the secondary and main buses. (i.e. no Ext AC power source is available)
Battery temperatures and charging rates should be continuously monitored on the ELECTRICAL page, after starting.
After the operating DC generator is connected to the related main bus (the start SELECT switch returns to the center position), it will help in the starting of the other engine, in parallel with the main and auxiliary batteries.
During the start process and 15 seconds following it, the DC GCU's are supplied with a 'CURRENT LIMIT' signal. This limits generator output. This is canceled when the engine START segment goes out. 3
After the APU generator is connected to the right main bus, it will help in the starting of the engines, in parallel with the main and auxiliary batteries as above.
A Ground Power Unit (GPU) can be connected to the DC external power receptacle, on the left side of the nose section and requires the Battery Master to be on. During an engine start the main and aux batteries are connected to the main buses and assist in the start.
After engine start, and the DC EXT PWR switch is set to OFF, the generators will come on line, (if the GEN switches are in the 1 or 2 position). Bus fault logic connects the main to the secondary buses, until the TRUs. come on line (ie have an AC power source)
The EPCU incorporates external DC power protection (22-31 VDC); an over/under voltage condition will cause the external ground power to disconnect. If the situation is rectified, moving the DC EXT PWR switch to OFF and then back to EXT PWR can reselect the external power source.
Electrical power sources for the AC portion of the EPGDS include:
-Two alternating current variable frequency generators
The power sources supply power to the following buses in order of priority:
-Left or Right AC bus
-Left or Right Galley bus
During normal mode of operation, each AC source supplies its dedicated bus. AC power is required for:
-Standby hydraulic pump
-Transformer Rectifier Units (supplements DC power)
-Auxiliary fuel pumps
Two 115 VAC generators mounted on the Propeller Reduction Gearbox (RGB), supply variable frequency AC power to the left and right AC buses. AC power sources are prevented from being operated in parallel.
AC generated power is available when the condition levers are in or above the MIN/850 position, and the GEN 1 and GEN 2 switches on the AC CONTROL panel are on.
If one AC generator fails, the associated #1 AC GEN or #2 AC GEN caution light comes on. An automatic cross tie function, controlled by the AC GCU logic circuits, ensures that all variable-frequency buses are powered, except galley power. If an excessive load is detected, (bus fault) the GCU isolates the bus, and turns on the appropriate L AC BUS or R AC BUS caution light. The associated AC GEN caution light may also be illuminated. Associated services supplied by the bus are no longer powered and their associated caution lights will be illuminated.
The # 1 AC GEN HOT or # 2 AC GEN HOT caution light comes on if an AC generator overheats.
The AC external power receptacle is in the right engine nacelle, or it can be installed on the right side of the nose.
An external power switch connects power directly to the left and right variable-frequency buses, supplying power to all AC and DC buses. During engine starts, the secondary buses are not connected to the main buses.
An AC PPU is installed on the right AC Contactor It monitors:
-voltage under or over
-frequency under or over
-Phase Rotation (A‑B‑C)
Circuit and current limiters protect electrical system power sources, component control circuits and bus distribution.
- flight compartment circuit breaker panels (4)
-DC Contactor box in the nose compartment
-AC contactor boxes in the left and right main landing gear wheel wells
-Two wardrobe circuit breaker panels
-Galleys circuit breaker panels
Each circuit breaker is identified by the:
- Identification label
- Alphanumeric location.
The fire protection system supplies detection, indication and extinguishing of fire or smoke conditions.
Indication and test functions are supplied for:
Lavatory smoke detection is only indicated in the lavatory and passenger compartment.
There are portable fire bottles in the flight and passenger compartments. Refer to Chapter 3 for a description of APU fire detection and extinguishing.
Advanced Pneumatic Detectors (APD) in the engine nacelles and one APD in the APU compartment provide overheat detection. Smoke detection is provided by two smoke detectors located in the aft baggage compartment, one in the forward baggage compartment and one in the lavatory. Indication is shown on the Fire Protection Panel (FPP), Caution and Waning Panel, and glareshield
The aircraft fire detection system monitors overheat conditions for engine fire zones. If this is sensed, the system supplies a visual and aural warning to the flight compartment.
Advanced Pneumatic Detectors (APDs)provide overheat detection for the 1) Primary Engine Zone (PEZ), 2) Leading Edge Zone (LEZ) and 3) Main Wheel Well (MW zone. The APDs also supply fault indications to the Fire Protection Panel.. The PEZ also includes the Propeller Electronic Controller (PEC).
The APDs use sensor tubes, filled with helium gas, sense overpressure (fire) and under-pressure (fault) The signals are processed by a control amplifier then sent to the fire protection panel.
It is possible to get a FAULT A indication and an engine fire warning at the same time.
During the fire detection test, the control amplifier is also tested. If the control amplifier fails, it will not cause complete loss of engine detection or extinguishing capability.
Two dual port fire bottles are installed forward and aft in the left wing root, for engine fire extinguishing. There are electrical connections for explosive squibs and the bottle monitor pressure switch.
Fire suppressant can be discharged into the left or right nacelles. The bottles are connected in a configuration that allows for up to two suppressant discharges into an engine nacelle. If the discharging the first bottle does not put out the fire, the second bottle can be discharged.
A BTL LOW amber advisory light, on the Fire Protection panel, comes on when a fire bottle is empty. Pulling the PULL FUEL/HYD OFF handle will show which bottle is empty the related FWD BTL or AFT BTL amber arming light will not come on as it should..
When an overheat condition occurs, the alarm signals are processed by the Control Amplifier. and
- Engine fire warning tone sounds
- Both ENGINE FIRE PRESS TO RESET switchlights (red) flash
- CHECK FIRE DET warning light (red) flashes
- Related PULL FUEL/HYD OFF handle light (red) comes on
Pushing either ENGINE FIRE PRESS TO RESET switchlight stops the engine fire warning tone and the flashing. These switchlights stay on steady until the fire overheat condition stops.
The fwd and aft bottle squibs are armed by pulling the PULL FUEL/HYD OFF handle. After arming, the extinguisher bottle is discharged by holding the EXTG switch to the FWD or AFT position. and the fire suppressant discharges into the engine zones.
Fire extinguishing for the baggage compartments is by two High Rate Discharge (HRD) fire bottles and one Low Rate Discharge (LRD) fire bottle. Each baggage compartment has one HRD fire bottle. The LRD fire bottle is shared between the FWD and AFT baggage compartments, and is located in the aft equipment bay.
The aft baggage compartment has two smoke detectors. One smoke detector is located in the rear and one in the front of the aft baggage compartment.
If one or both aft baggage compartment smoke detector senses smoke, the:
- SMOKE warning light (red) ‑ flashes
- MASTER WARNING light (red) ‑ flashes (tone)
- AFT SMOKE segment (red)and EXTG segment (white)
- VENT VALVE INLT light (white)
- VENT VALVE OTLT light (white)
- AFT ARM segment (amber)
Pushing the SMOKE/EXTG switchlight discharges the HRD fire suppressant into the aft baggage compartment. The AFT ARM light goes off and the AFT LOW light comes on.
After a seven minute delay, the LRD fire bottle automatically discharges into the aft baggage compartment. The seven minute delay is to maximize the amount of suppressant in the baggage compartment. The FWD LOW light comes when the LRD bottle pressure is low.
The forward baggage compartment has one smoke detector.
If the fwd baggage compartment smoke detector senses smoke, the following lights illuminate:
- SMOKE warning light (red) ‑ flashes
- MASTER WARNING light (red) ‑ flashes (tone)
- FWD SMOKE segment (red) and EXTG segment (white)
- FWD ARM segment (amber)
Pushing the SMOKE/EXTG switchlight discharges the forward HRD fire suppressant into the forward baggage area. The LRD fire extinguisher bottle will discharge at the same time.
The FIRE BOTTLE FWD ARM light will go off and the FWD LOW light will come on (loss of bottle pressure). The AFT LOW light comes on when the LRD bottle pressure is low.
There are four portable fire extinguishers, one in the flight compartment, and three in the passenger compartment. A gauge on each extinguisher shows the serviceable range (Green), overcharge range (Yellow), and recharge range (Red). Each extinguisher contains Halon 1211, which is effective on electrical, oil and fuel fires. The extinguishant is not corrosive or toxic, and will not freeze or cause cold burns. A red safety catch prevents accidental trigger movement and discharge.
There is one smoke detector in the lavatory compartment (Figure 8‑8).
If the smoke detector senses smoke, the:
- Detector audible warning tone sounds
- Detector alarm light (red) comes on
- LAV SMOKE light (red), on all 3 Advisory Light Panels (ALP), come on
- Chime (high) sounds in the cabin speakers
There is no lavatory smoke indication in the flight compartment.
The lavatory smoke detector is tested by pushing a self-test pushbutton on the detector. Pushing the self-test pushbutton simulates a smoke condition, and causes the same indications. If the interrupt pushbutton is pushed during a smoke test, the audible warning tone and chime are silenced, and the LAV SMOKE lights go off.
The waste bin, in the lavatory compartment, is protected by a thermally activated fire bottle (Potty Bottle) with no electrical interface. The Potty Bottle has dual discharge tubes. If a fire occurs in the waste bin, the temperature of the end caps of the tubes increases. Once the temperature increases to a set point, the fusible seals melt and release the end caps from the discharge tubes. The suppressant is then discharged into the waste bin.
The Dash 8 Q400 primary flight controls consist of rudders, ailerons and elevators. Spoilers assist the ailerons for roll control. Secondary flight controls consist of flaps.
All flight controls may be operated from either the pilot or copilot's seat. The rudders provide yaw control, the ailerons and spoilers roll control, and elevators pitch control .The rudder, spoilers and elevators are hydraulically powered, and designated the Powered Flight Control Surfaces (PFCS). A gust lock system is provided for the aileron controls, to protect the ailerons from damage due to strong wind gusts.
The spoilers assist the ailerons in providing roll control, and reduce lift after the aircraft touches down.
PFCS positions are shown in the Permanent Systems Area (PSA) of MultiFunction Display 1 (MFDI). These PFCS positions are transmitted to the PSA from the control surfaces. Trim indicators show trim position of the flight controls. Advisory lights indicate system operation, and caution lights indicate flight control malfunctions.
A hydraulically powered rudder provides yaw control. The rudder pedals control the rudder. The pilot's and copilot's rudder pedals are connected to each other through an interconnect rod. A mechanical feel and trim unit, provides simulated aerodynamic forces on the rudder pedals during flight. A yaw damper operates through the feel and trim system to improve directional control.
The rudder control system provides directional control of the aircraft. The rudder consists of the fore rudder and trailing rudder.
Rudder position can be monitored in the Permanent System Area (PSA) of MFD1
The fore rudder is attached to the vertical stabilizer and operated by two Power Control Units (PCUs). The PCUs are installed one above the other at the midpoint of the vertical stabilizer. No. 1 hydraulic system powers the lower PCU and No. 2 hydraulic system powers the upper PCU. Moving the rudder pedals operates both PCUs. If either hydraulic system fails, the remaining PCU provides rudder control.
The trailing rudder is attached to the fore rudder by push rods and deflects mechanically with movement of the fore rudder. The trailing rudder deflects twice as far as the fore rudder.
A rudder input restrictor mechanism, limits rudder pedal travel with flap selector lever operation. The flap selector lever is mechanically linked to the copilot's rudder control. With FLAPS lever set at 0, pushing either rudder pedal to the stops, deflects the fore rudder 12º left or right of center. With FLAPS lever set at 5º or more, pushing either rudder pedal to the stops, deflects the fore rudder 18º left or right of center.
Hydraulic pressure supplied to both PCUs is regulated by the Flight Control Electronic Control Unit (FCECU) as airspeeds vary.. As airspeed increases, FCECU reduces the hydraulic pressure available to the PCUs to reduce rudder deflection. Airspeed information is supplied by the Air Data Units (ADUs).
Rudder pedal adjustments are provided for both sets of rudder pedals. A cable connecting the pilot's and copilot's brake pedals, allows for operation of the brake system from either pilot position.
The rudder feel trim and summing unit, provides artificial feedback forces to the rudder pedals. This simulates aerodynamic forces from the rudder during flight.
Inputs from the rudder pedals and yaw damper are applied to the summing unit. The unit and then transmits the resultant command, as a single input, to the rudder PCUs.
Turning the RUDDER trim selector, located on the center console, operates an electrical trim actuator to reposition the rudder feel unit neutral point. The rudder pedals also move when trim is selected. Trim indication is shown on the RUDDER trim indicator.
Turning the RUDDER trim selector to the first graduation, produces a slow trim rate. Turning the selector fully to the second graduation produces a fast trim rate. system is powered from the Left Essential bus through two circuit breakers:
RUD TRIM ACT- F7 for the trim actuator
RUD TRIM IND - G7 for the RUDDER trim indicator.
If the rudder trim switch fails, causing uncontrolled rudder trim; a limit switch shuts off the electrical power to the trim actuator. There is mechanically stopped at the maximum trim setting
Trim actuator position is shown on the RUDDER trim indicator. If the trim signal fails, the trim actuator remains functional but an off scale deflection is shown on the RUDDER trim indicator
The yaw damper actuator supplies automatic compensation for minor yaw acceleration during flight. It also improves directional stability and provides turn coordination. Yaw damper authority is 4.5º maximum of rudder deflection either side of center. The yaw damper gets its inputs from Flight Guidance Modules No. 1 and No. 2 and needs both inputs for operation.
If a jam occurs in either rudder PCU, is indicated by the RUD 1 or RUD 2 PUSH OFF switchlight on the glareshield. RUD 1 or RUD 2 switchlight must be pushed to depressurize the affected PCU. The OFF segment stays on as a reminder that the switchlight has been pushed off and #1 RUD HYD or #2 RUD HYD caution light will come on.
The FCECU will now modify the hydraulic pressure to the operative PCU to maintain the required rudder authority as airspeed varies.
Ailerons assisted by spoilers provide roll control. The aileron control system and flight spoiler control system are independent systems. Both systems are mechanically interconnected to allow simultaneous operation for normal roll control. The AFCS can provide input commands to the roll control system.
• Each wing has one aileron and two flight spoilers
• The pilot's wheel is connected directly to the flight spoilers
• The copilot's wheel is connected directly to the ailerons
• Ailerons are mechanically controlled and cable operated
• Flight spoilers are mechanically controlled and hydraulically powered
If a roll control jam occurs, the spoiler control system can be separated from the aileron control system. The pilot with the unjammed control handwheel will have roll control.
An aileron is located on the outboard trailing edge of each wing. Control is conventional by the control wheels.
Each aileron has a geared tab. When the ailerons are deflected up or down, its geared tab moves in the opposite direction. This provides aerodynamic assistance to the pilot.
A ground adjustable trim tab is installed on the right hand aileron.
The three position AILERON trim switch controls the trim actuator. The switch is spring-loaded and returns to the center off position.
The aileron trim system is electrically powered from the Left Essential bus through circuit breakers:
AIL TRIM ACT - G8 for the aileron trim actuator
AIL TRIM IND - H8 for the aileron trim indicator
The amount of trim is shown on the aileron trim indicator.
The Aileron Trim and Centering Unit (ATCU) is a spring that operates between the trim actuator and the forward aileron quadrant. It provides aileron trim input as a spring bias and automatic centering of the ailerons so that when the pilot control wheel is released, the wheel returns to the neutral position set by the ATCU.
Trimming deflects both ailerons and repositions the neutral position of the control wheel.
When the autopilot is engaged, a MISTRIM [TRIM L WING DN or R WING DN] message on the PFD indicates to the pilot the AILERON trim input required, to remove control forces. When accomplished the message is removed.
If the aileron trim switch fails trimming is electrically and mechanically stopped at the maximum trim setting.
Each wing has an inboard and outboard roll spoiler. The roll spoilers operate with the ailerons to assist roll control of the aircraft. The roll spoilers are extended and retracted hydraulically.
Pushing either SPLR1 or SPLR2 switchlight, cuts hydraulic pressure to its related spoiler extend port The associated ROLL SPLR INBD HYD or ROLL SPLR OUTBD HYD caution light will come on
There are three modes of spoiler operation:
The spoilers operate in proportion to the up going aileron to provide roll control. The No. I hydraulic system powers the inboard spoilers and No. 2 hydraulic system powers the outboard spoilers. At airspeeds more than 165-170 KIAS, the FCECU disables the outboard spoilers.
If the outboard spoilers are not shutting on/off correctly the SPLR OUTBD caution light comes on.
There are two electrically operated lift-dump valves in the each spoiler system for ground spoiler operations. They are hydraulically connected in series; both valves must open together before the spoilers can extend. When the lift-dump valves are energized open, spoilers extend.
The lift-dump valves are energized by signals from the FCECU. Spoilers extend on touchdown when:
- The FLIGHT/TAXI switch is in the FLIGHT position.
- Both POWER Levers are less than FLIGHT IDLE + 12º.
- Main Landing Gear Weight-On-Wheels (WOW) proximity sensors detect the aircraft is on the ground, and the Proximity Sensor Electronic Unit (PSEU) WOW equations are satisfied.
Then the roll spoilers extend. and the ROLL INBD and ROLL OUTBD advisory lights come on. This decreases the lift on the wings to assist in maximum braking efficiency.
If the paired lift-dump valves are not in the same position, The related ROLL SPLR INBD GND or ROLL SPLR OUTBD GND caution light will come on.
The FLIGHT/TAXI switch is spring-loaded to FLIGHT position and must be manually set to the TAXI position. It is kept in the TAXI position by a magnetic latch. When on ground, all spoiler panels can be retracted by setting the FLIGHT/TAXI switch to TAXI.
When both POWER Levers are moved to a position more than FLIGHT IDLE + 12º, (i.e. in take-off) the latch is de-energized, and the switch moves to the FLIGHT position.
If there is a malfunction in the spoilers or aileron control system and a roll control jam occurs a roll disconnect system is provided this consists of a clutch between the pilot's and copilot's control columns.
The ROLL DISC handle is pulled out to the limit and turned 90º clockwise or counterclockwise. This disengages the clutch, and isolates the jammed system from the operating system. The pilot with the unjammed wheel will have roll control.
Left Control Wheel Free
If the left control wheel is free, only roll spoilers will operate Roll control forces will be low and the tendency to overcontrol should be avoided.
Right Control Wheel Free
If the right control wheel is free, only ailerons will be operational. Roll control will be reduced and forces will be normal.
If the control wheel is rotated more than 50º from neutral to maintain wings level, SPLR 1 and SPLR 2 switchlights will come on This may be due to one or both roll spoilers on the same side being extended.
If the SPLR 1 and/or SPLR 2 switchlights stay on continuously, they must be pushed off to depressurize the PCU(s) and retract the related spoiler(s). The OFF segment stays on to indicate the switchlights have been pushed off. This will turn on the ROLL SPLR INBD HYD and/or ROLL SPLR OUTBD HYD caution lights. The ROLL SPLROUTBD HYD caution light will not come on until speed is less than 165 KIAS. Roll spoiler positions may be monitored on the PSA of MFD1.
Two mechanically controlled and hydraulically powered elevators maintain pitch control of the aircraft. 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. Both control columns are connected to each other by the pitch disconnect system so that they both operate together.
Fore and aft movement of both control columns is transferred, through two fully independent cable and pulley control circuits, to the elevator PCUs.
There are three identical hydraulic PCUs (hydraulic Power Control Units) on each elevator. The No.1 hydraulic system supplies power to the outboard PCUs. The No.2 hydraulic system supplies power to the center PCUs. The standby No.3 hydraulic system supplies power to the inboard standby PCUs when required.
Pushing the guarded HYD #3 ISOL VLV pushbutton, on the HYDRAULIC CONTROL panel, manually opens the #3 isolation valve, pressurizing the inboard PCUs. This causes the ELEVATOR PRESS caution light to come on if the No. 1 and No. 2 hydraulic systems are also operating. The #3 isolation valve will also open automatically if there is a failure on either main unit
Two pitch trim actuators do pitch trim. The actuators are controlled automatically by the autopilot, or manually by the trim switches on either control wheel. Elevator trim position is shown on the elevator trim indicator located on the left side of the center console.
If a mismatch occurs between the left and right elevator, the ELEVATOR ASYMMETRY caution light comes on. Elevator position indication is shown on the PSA of MFD1. Gust protection for the elevators is supplied by trapped hydraulic fluid in the actuators, when the system is depressurized.
A left and right Pitch Feel and Trim Units (PFTUs) provide artificial pitch feel. The PFTUs are installed in the vertical stabilizer. The right PFTU controls the right elevator and the left PFTU controls the left elevator. Pitch commands from the control columns are transferred to the elevator PCUs that move the elevators. Thus, actual aerodynamic forces are not felt at the control columns.
Centering springs in the PFTU systems, help to return the elevators to the neutral position. Two pitch trim actuators installed on top of the PFTUs supply elevator trim.
Both pitch feel actuators operate at the same. As airspeed varies, the FCECU commands the pitch feel actuators to supply the correct artificial forces to the control columns. The elevator column force increases with column displacement as a function of airspeed and normal acceleration of the aircraft. Air Data Units (ADUs) supply airspeed information to the FCECU.
If one pitch feel actuator fails, the other actuator will operate normally. The ELEVATOR FEEL caution light will come on. The elevator control system continues to operate but with reduced artificial feel.
Pitch trim is accomplished by two pitch trim actuators to trim the elevators. The elevator trim actuator is controlled electrically by the trim switches on either control wheel, or automatically by the autopilot.
Pitch trim signals from are prioritized by the FCECU in the order: pilot, copilot and autopilot.
The FCECU controls the elevator pitch trim rate according to the airspeed inputs from the ADU of the aircraft. The FCECU adjusts the trim rate between 150 KIAS and below (high-speed mode) to 250 KIAS and above (low speed mode)
Elevator trim control is provided through the actuation of trim switches located on the outboard handgrip of each control wheel. The pitch trim switches are divided into two halves. Both halves must be operated for pitch trim commands. They are thumb-operated switches, which are spring loaded to the center off from NOSE DN and NOSE UP positions.
When the switches are pushed forward to NOSE DN position, a nose‑down trim is commanded. When the switches are pulled aft to NOSE UP position, a nose‑up trim is commanded. If FCECU detects a pitch trim command for longer than 3 seconds, an aural warning will sound and the ELEVATOR TRIM SHUT OFF switchlight on the glareshield panel will come on.
The aural warning will stop and the ELEVATOR TRIM SHUT OFF switchlight will go off when the pitch trim command is removed or the switchlight is pushed. If the left or right ELEVATOR TRIM SHUT OFF switchlight is pushed; the elevator trim is turned off.
The elevator trim indicator is located on left side of the center console, and shows the elevator trim position.
The indicator is labeled NU for nose up, ND for nose down and TO for
takeoff. A white band next to the TO label shows the takeoff trim range. If the
elevator trim is:
weight on the nose wheel and
a takeoff aural warning will sound.
Nose full down =circuit depowered
Nose full up = trim selected off
The elevators are automatically retrimmed whenever the flaps are moving between 15º and 35º. Nose down pitch trim is commanded when flaps are extended, and nose up pitch trim is commanded when flaps are retracted. The Flap Auto Trim activates and deactivates automatically. Function can be verified on the trim indicator.
Flap Auto Pitch Trim is active when:
- Flaps selected from 15° to 35º, and
- The autopilot is not engaged, and ???
- The airspeed is less than 180 KIAS, and
- Manual pitch trim is not commanded.
Flap Auto Trim will temporarily disengage if manual pitch trim is applied.
Flap Auto Pitch Trim will automatically disengage, if the above are exceeded or when:
- The aircraft is on the ground (WOW), or
- Commands are in excess of the pitch limits, or
-failures within the AFCS or Flight control system occur.
The left and right control columns are mechanically connected to each other through the pitch disconnect mechanism (a clutch). If a pitch jam occurs in either control circuit, the two control columns can be disconnected from each other by using the pitch disconnect handle located on the left side of the center console. When the handle is pulled out and rotated 90º the clutch disengages and disconnects the two control columns from each other. The pilot with the free control column will have pitch control.
Two single‑slotted inboard and outboard fowler flaps are attached to the trailing edge of each wing The flaps are connected to screw jacks that are operated by a primary drive shaft. A Flap Power Unit (FPU) operated by the flap selector, operates the flap drive system and moves the flaps to their selected positions. The flaps are electronically controlled by the Flap Control Unit (FCU) and powered by No. 1 hydraulic system.
The flap quadrant has gates at five positions corresponding to the five possible flap positions; 0°, 5°, 10°, 15°; and 35°. These Flap positions are shown on the FLAP indicator.
The Flap Control Unit (FCU) monitors and controls flap movement. It also controls the automatic operation of the Standby Hydraulic Pump and the PTU. If there is a flap fault, the FCU causes either the FLAP POWER or FLAP DRIVE caution light to come on.
The FLAPS selector lever moves in the five gates of the flap quadrant. A trigger,(and flap arming switch) located below the FLAPS selector lever, must be pulled up before the lever can be moved from one gate to the next. The trigger must be released in the desired gate before the flaps start moving.
The FLAPS selector lever is mechanically connected to the rudder restrictor on the right rudder forward quadrant. This restricts rudder pedal travel to 12° left or right of center, if the FLAPS selector lever is set to 0°. The FLAPS selector lever at the other gates does not restrict full rudder travel.
The Flap Power Unit (FPU) converts The No. 1 hydraulic system power to rotary mechanical power to operate the flap actuators. The FPU receives flap actuation signals from the FCU. It also sends feedback signals to let the FCU monitor flap movement.
When the FLAPS selector lever is in the 0° position, all flaps are retracted, and the flap position indicator points to 0°. Movement of the FLAPS selector lever requires raising the arming switch that sends an enabling signal to the FCU. When the FCU receives the enabling signal, it commands the FPU to start moving the flaps to the new position. When the selected flap position‑ has been reached, the FCU stops any further flap movement. The flap actuators lock the flaps in the selected position.
Electronic monitoring of the flaps minimizes non‑selected flap movement. It also minimizes flap asymmetry, if the flaps drive line fails.
There are four flap actuators, two per flap, installed on each wing. The flap actuators move the flaps to a selected position, and lock them in the selected position against aerodynamic forces.
The Flap Position Indication Unit (FPIU)
-supplies indication of the current flap position to the FLAP indicator flap position data to the flight compartment and to other aircraft systems for indication and monitoring purposes.
-monitors flaps position during flap deployment and signals position data to the FCECU, IFC and ADU.
The aileron gust lock system mechanically locks the ailerons in the neutral position. A CONTROL LOCK lever on the power quadrant, forward of the POWER levers operates the gust lock system. When the CONTROL LOCK lever is in the ON position, the POWER levers will move only as far as the CONTROL LOCK lever.
The CONTROL LOCK lever is spring loaded to the forward OFF position. A trigger under the lever locks it in the aft ON position, with the copilot's control wheel in the neutral position. If a gust lock cable fails, the gust lock mechanism fails safe to the unlocked position.
To undock the aileron gust lock, the CONTROL LOCK lever must be pulled aft, and the trigger under the gust lock lever squeezed to release it. The CONTROL LOCK lever can then be moved forward to the OFF position.
The Rudder and the Elevator Control Systems are hydraulically powered flight control systems. Gust lock protection for these two systems is provided by retained hydraulic pressure in the flight control actuators. The retained pressure provides stiffness to the flight control surface, which resists movement from wind forces.
Fuel is contained in two integral main wing tanks. The fuel system provides for indicating, storing, venting, fuel feeding and scavenging, refueling/defueling, and transferring. Only tank to tank transfer is available; there is no engine crossfeed capability
A fuel gauging system supplies and displays quantity data in the flight compartment and the refuel-defuel panel. Fuel quantity may also be checked on the ground by use of the magnetic dipsticks. Each wing tank includes a surge bay and a collector bay. The left tank supplies fuel to the left engine and the optional Auxiliary Power Unit (APU). The right tank supplies fuel to the right engine. A vent system keeps the air pressure in the fuel tanks between structural limits. Fuel can be transferred between the tanks for lateral balancing or for fuel management. A single point pressure refuel/defuel system shares selected common components with the fuel transfer system. Gravity refueling is done through two overwing fuel filler points.
Fuel quantity and temperature data is shown in white on the Engine Display (ED) and the FUEL Page. White dashes replace the digits or simulated dials if the valid data is not available.
A Fuel Quantity Computer (FQC) uses nine capacitance type fuel probes in each tank to determine the total fuel quantity. The fuel quantity of each tank is shown in digital form on the bottom center of the Engine Display (ED).
The FUEL Page also shows the fuel quantity of each tank on two simulated analog dials and below that the total fuel quantity in digital form. The fuel quantity can be shown in kilograms (KG) or pounds (LBS). The fuel flow for each engine is shown in digital form on the ED in units of KG/H or PPH.
The fuel temperature (°C) in the left collector tank is shown in digital form on the FUEL Page, just below the left fuel quantity dial. The engine inlet temperature of the fuel, after it has passed through the Fuel/Oil Heat Exchanger (FOHE) for the left and right engines, is shown in digital form on the bottom of the ED, just below the related fuel tank quantity. The digits are shown in white with a ± sign, and change to yellow or red if the fuel temperature is above or below limits. A °C unit is shown between the two inlet temperatures.
The fuel quantity of each tank is also shown on the refuel/defuel panel located at the back underside of the right nacelle. One (??)magnetic dipstick, on the underside of each tank, can also be used to give an independent mechanical indication of the fuel quantity in litres or U.S. gallons. The magnetic dipstick is a calibrated rod with a magnet attached to the top. It moves within a tube that extends vertically from the bottom of the fuel tank. A float moves up and down on the outside surface of this tube to match the fuel level in the tank. The float contains a magnet that is attracted to the magnet on the magnetic dipstick, holding the dipstick at that level. The magnetic dipsticks are not accurate when the fuel quantity is near full or empty.
There are two integral (wet) wing tanks that extend laterally from the fuselage to the rib just inboard of the ailerons. Each wing tank is divided into three sections:
- Surge bay
- Main tank
- Collector bay
The surge bay is located between the two ribs inboard of the aileron. The main tank extends from the surge bay to the fuselage and collector bay. The collector bay is located at the inboard and aft part of the wing tank. Fuel is contained in the main tanks and the collector bays. Total useable fuel from both tanks is 5,318 kg (11,724 Ibs.). The maximum lateral imbalance permitted between tanks is 272 kg (600 Ibs.). Water drain valves are located at the low points of the tanks, surge and collector bays.
The surge bay is used for main tank venting and fuel recovery. Two outboard float vent valves and one inboard vent line, control the pressure between the related surge bay and main tank The two float vent valves, located near the top of the tank, open and close depending on the fuel level in the main tank. Each surge bay is vented through integral standpipes to two separate NACA vents on the bottom of the wings. During flight, any fuel that may spill into the surge bay, is returned to the tank by the reduced pressure in the main tank, as fuel is used.
The collector bay supplies engine fuel regardless of aircraft attitude. Fuel scavenge ejector pumps pull fuel from tank low points, and pump it into the collector bay. A primary ejector pump in the collector bay then supplies fuel at a constant low‑pressure to the engine. High‑pressure motive flow is used to operate the scavenge and primary ejector pumps.
Flapper valves are located at the base of the collector bay and wing ribs. The flapper valves only let fuel flow inboard, making sure the collector bay always has fuel.
The #1 TANK or #2 TANK FUEL LOW caution light comes on if:
• park brake is off
• related collector bay level drops below approximately 150 kg (305 Ibs.)
• related engine is operating
Fuel to each engine is fed from the collector tank, from a primary ejector pump or an AC (Variable Frequency) driven auxiliary pump, and delivered to the engine driven pump. If the engine driven pump inlet pressure drops below limits, the related #1 or #2 ENG FUEL PRESS caution light comes on.
The auxiliary pump in each collector bay serves as a back up source of fuel boost pressure for takeoff and landing,
and in case the related primary ejector pump does not supply the necessary fuel pressure. Related TANK1 or TANK 2 AUX PUMP switchlights on the FUEL CONTROL TRANSFER panel control the auxiliary pumps manually A TANK1 orTANK 2 AUX PUMP switch indicator, on the FUEL Page, shows the position of the switchlight. When the pump is supplying sufficient fuel pressure, the TANK1 or TANK 2 AUXPUMP 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, is pulled . White and green FUEL advisory lights, on the Fire Protection Panel (FPP), show the valve position.
The fuel is filtered and heated by the Fuel Oil Heat Exchanger (FOHE) a before entering the engine pump. If the fuel
filter becomes blocked, fuel bypasses the filter. If a related fuel bypass is impending, the #1 or #2 FUEL
FLTR BYPASS caution light comes on.
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 Ibs.), a yellow [BALANCE] message flashes just above the FUEL legend of the ED. The message keeps flashing until the imbalance is corrected. The fuel quantity digital values also change to yellow during a fuel imbalance. An imbalance condition is also 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. When the TRANSFER switch is selected, the auxiliary fuel pump in the donor tank automatically operates to pump fuel to the receiver tank. A fuel pressure signal from the operating pump side causes the related ON switchlight segment to turn green. Fuel transfer indications are also shown on the FUEL Page. Electrically operated fuel transfer shutoff valves open for fuel transfer and close when the transfer is stopped. The level control shut off valve of the receiver tank opens for fuel transfer.
Once selected, fuel transfer will continue until deselected, or until the high level sensor in the receiver tank detects an overfill condition..
Fuel cannot be transferred if the FUELING ON caution light is on. The FUELING ON caution light indicates the refuel/defuel access door is open.
The refuel/defuel panels controls all refuel and defuel operations. Access to the panel is through a flush door on the rear underside of the No. 2 nacelle. The FUELING ON caution light will come on when the refuel/defuel door is open with electrical power supplied. DC power must be available for refueling. Refueling can be done either automatically (PRESELECTREFUEL), or manually (REFUEL).
When the rotary selector is turned to the PRESELECT REFUEL or REFUEL position, the MASTER VALVE CLOSED light goes out to show that the refuel/defuel shutoff valve is open. The refuel/defuel shutoff valve will close when the refuel/defuel door is closed, regardless of the rotary selector position. The PRECHECK, OPEN, CLOSE switches must set to OPEN for refueling. During pressure refueling, fuel pressure opens a vent/dump valve in each tank to vent the tanks through the surge bay.
For automatic refueling, the desired quantity is selected adjacent the PRESEL display on the Refuel/Defuel Indicator (RDI), by using the INCR DECR toggle switch. When the selected quantity is reached, the level control shutoff valve automatically stops fuel flow to the related tank.
During manual refueling, the desired quantity in each tank is controlled by the PRECHECK, OPEN, CLOSE switches. The RDI will display the individual tank quantities. With the rotary selector turned to the REFUEL position, refueling continues until the PRECHECK, OPEN, CLOSE switches are in the CLOSE position, the selector switch is turned to the OFF position, or the high level sensors sense a full tank condition.
When fueling is in progress, advisory lights on the refuel/defuel control panel indicate the position of the fuel control valves. The amber DUMP VALVE OPEN indicator light comes on when the related vent/dump valve opens. The vent/dump valve allows air to vent from the tank as it fills, and prevents damage should the tank overfill due to failure of both normal and backup shut‑off features. In the latter case, fuel flows into the surge bay through the refuel vent valve and, if it reaches the height of the standpipes, is spilled overboard through the surge bay NACA vents.
Two PRECHECK, OPEN, CLOSE switches test the automatic shutoff operation of the related high level control unit during refueling, ensuring operation of the overfill shut off system. With the switch in the PRECHECK position a full tank is simulated for the related tank, and the high-level sensor shuts off refueling to that tank by closing the level control shut off valve. This causes the related amber REFUELSHUTOFF light to come on. Refueling restarts when the switch is released. If DC power is stopped during refueling, refueling stops.
Gravity refueling can be done through the wing mounted gravity refuel adapter on the top of the wing .
When the rotary selector is turned to the PRESELECT DEFUEL position, the FGC closes the refuel/defuel valves to automatically stop defueling at the preset quantity. When the rotary selector is turned to the DEFUEL position, defueling continues until the selector switch is turned to the OFF position. If AC (variable frequency) power is available, the related auxiliary fuel pump operates to help the defueling process. Without an AC power source, suction defueling is required. DC power is required for defueling regardless of AC power availability.
The maximum refueling pressure is 50 psi, which gives a refueling flow rate of approximately 125 gallons per minute. A minimum of 20-psi refueling pressure is necessary for the system to operate properly.
The Dash 8 Q400 has four hydraulic systems, two main systems, an auxiliary system, and an alternate landing gear extension system.
The No. 1 and No. 2 main systems power the:
• Landing Gear
• Nose Wheel Steering
• Rudder and Elevators
The No. 3 auxiliary hydraulic system supplies backup pressure to the elevators. The alternate landing gear extension hydraulic system is used as a backup for main landing gear extension.
Hydraulic power is by three hydraulic systems, No. 1 (left), No. 2 (right) and No. 3 (aft). The No. 1 and No. 2 hydraulic systems are normally pressurized by the related Engine Driven Pump (EDP). A Standby Hydraulic Pump is a backup power for the No. 1 hydraulic system. A Power Transfer Unit (PTU) is a backup for the No. 2 hydraulic system.
The No. 3 hydraulic system is powered by a DC Motor Pump (DCMP) The No. 3 hydraulic system is backup power for the elevators.
The alternate landing gear extension hydraulic system is used to extend the main landing gear, if required.
NO. 1 AND NO. 2 HYDRAULIC SYSTEMS SYSTEM OPERATION
The No. 1 and No. 2 hydraulic system supplies hydraulic fluid, from the pressurized reservoir through the Firewall Shutoff Valve to the EDP. Pressurized fluid is then available to subsystems at a nominal pressure of 3000 psi., before returning to the reservoir. If an EDP fails, the related #1 or #2 ENG HYD PUMP caution light comes on.
The No. 1 system powers the:
• Anti-Skid and Normal Brakes
• Inboard Roll Spoilers
• Rudder (Lower PCU)
• Elevators (Outboard PCUs)
The No. 2 system powers the:
• Landing Gear
• Outboard Roll Spoilers
• Nose Wheel Steering
• Emergency/Parking Brakes
• Rudder (Upper PCU)
• Elevators (Center PCUs)
The hydraulic system reservoirs store / supply fluid to the hydraulic systems. The No. 1 hydraulic system reservoir (8 U.S. Quarts) is installed in the No. 1 nacelle. The No. 2 reservoir (12 U.S. Quarts) is installed in the No. 2 nacelle. The reservoir uses system output pressure to pressurize itself and provide supply pressure to the EDPs.
If the Hydraulic fluid in each reservoir is above temperature limits, the related #1 or #2 HYD FLUID HOT caution light comes on.
The No. I and No. 2 hydraulic systems each have an isolation valve which is normally open. If the system fluid quantity is low, the isolation valve closes and the #1 or #2 HYD ISO VLV caution light comes on. Hydraulic power is no longer available for:
• No. I system:
Normal brakes and inboard spoilers
• No. 2 system:
Gear, spoilers, nose wheel steering, and emerg brakes
The No. I and No. 2 hydraulic systems each have a firewall shut-off valve which, when closed, stops the flow of hydraulic fluid to the EDPs. The battery bus powers the hydraulic shut off valves. Each firewall shut-off valve has two HYD advisory lights on the Fire Protection Panel. The green advisory light indicates valve OPEN. The white advisory light for valve CLOSED.
The Firewall Shutoff Valve will close if:
• ENGINE I or ENGINE 2 PULL FUEL/HYD OFF handle is pulled
• No. I or No. 2 Hydraulic reservoir fluid quantity is low
• Hydraulic fluid over temperature condition has occurred.
The No. I and No. 2 hydraulic systems each have a hydraulic fluid to fuel heat exchanger, in each fuel collector bay, to cool the related hydraulic fluid.. A bypass valve controls the flow of hydraulic fluid to and from the heat exchanger.
The Standby Hydraulic Pump is a backup source of power for the No. I hydraulic system. The Standby Hydraulic Pump flow rate is less than the EDP; thus services will operate at a slower rate. It is installed in the No. I engine nacelle. The Standby Hydraulic Pump is driven byan AC motor powered by the right 115 VAC bus, with backup power from the left 115 VAC bus.
When the STBY HYD PRESS switchlight is pushed, the Standby Hydraulic Pump is energized, and the green ON legend comes on indicating the switchlight is pressed. Standby pump pressure is shown on the STBY HYD PRESS indicator.
The Standby Hydraulic Pump automatically operates if AC power is available:
• Park brake is off, and
• Flaps are selected greater than 0°, and
• No. I reservoir is not empty OR
• Airborne (WOW) and
• No. I engine fails
If the Standby Hydraulic Pump overheats, the #I STBY HYD PUMP HOT caution light comes on.
A priority valve is included in the No. I hydraulic system and is normally open. If the hydraulic pressure in No. I system decreases below 2100 psi, the priority valve closes. This shuts off hydraulic power to the flaps and PTU. Hydraulic pressure is maintained to the elevators, rudder, inboard spoilers and brakes.
A PTU is a backup source of hydraulic pressure for the No. 2 hydraulic system. The PTU uses hydraulic pressure from the No. 1 system to power a hydraulic motor. The motor then operates a hydraulic pump to pressurize the No. 2 system. Hydraulic fluid is not shared or transferred between No. 1 and No. 2 hydraulic systems during PTU operation. Hydraulic fluid must be available in the No. 1 and No. 2 system for PTU operation.
The PTU may be selected on manually or automatically. Manual selection is by pushing the PTU CNTRL switchlight. The green ON legend indicates the PTU is supplying hydraulic pressure.
The PTU automatically operates if:
No. 2 reservoir is not empty, and
Park brake is off, and
Flaps are selected greater than 0, and
No. 1 EDP pressure is more than 2400 psi
The No. 3 hydraulic system is an independent system which powers the left and right inboard elevator PCUs. The system operates automatically if the No. 1 and/or No. 2 hydraulic systems fail, or if a dual engine failure occurs. The No. 3 hydraulic system can also be engaged manually by pushing the HYD #3 ISOL VLV switchlight. Once pushed, an amber OPEN legend on the switchlight will come on. An accumulator and an isolation valve are also installed in the No. 3 hydraulic system.
A 28-volt 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 No. 3 system reservoir is 2.6 quarts.
The DCMP operates intermittently and is controlled by two 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 go 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.
The isolation valve isolates the elevators from No. 3 hydraulic system pressure. During normal flight operation, the accumulator isolation valve is energized closed. 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 and indicating switchlight position.
An additional pressure switch is installed downstream of the isolation valve. It turns on the ELEVATOR PRESS caution light if No. I, No. 2, and No. 3 hydraulic systems are supplying pressure to all six elevator actuators.
Hydraulic system pressure and reservoir quantity information is shown on the Permanent System Area (PSA) of MFD2. If either MFD malfunctions or is off, the data is shown in digital composite format on the other MFD.
System malfunctions are shown on the caution and warning panel.
The indications are:
• Park Brake pressure
• Standby system pressure
• No. 1, No. 2 and No. 3 system pressure
• No. 1, No. 2 and No. 3 quantity
The alternate landing gear extension system is used, if the main gear does not free fall to a locked position, during an alternate gear extension. The system supplies hydraulic power to downlock the main landing gear.
The Alternate Landing Gear selector valve is located below the flight compartment floor, and is normally in the open position. Opening the Landing Gear Alternate Extension door fully, closes the MLG selector valve, and gives access to the hand pump socket. Operating the hand pump lever provides pressure to the alternate landing gear actuators. This hydraulic pressure downlocks the main gear, if it did not freefall into position during an alternate extension.
A hand pump socket and lever operate the alternate landing gear extension system. The hand pump socket is located below the Landing Gear Alternate Extension door, in the flight compartment floor, adjacent to the copilot's seat. The hand pump lever, behind the copilot's seat, must be inserted into the socket to operate the extension system, after isolation of the No. 2 hydraulic system. The pump draws hydraulic fluid from an alternate landing gear system reservoir.
The alternate landing gear system reservoir is located in the nose compartment of the aircraft. The reservoir supplies the hydraulic fluid for the alternate landing gear extension system. The reservoir capacity is about 1 U.S. Quart.
Min Quantity #1, 2,and 3 systems - 50% subject to associated Temp. chart.
The Dash S Series Q400 Aircraft is approved for flight into known icing conditions. Ice and rain protection includes deicing, anti-icing, rain removal systems. And an ice detection system.
The ice detection system uses probes to actively check for icing conditions, and give indications in the flight compartment.
The deicing system uses engine bleed air to operate inflatable boot sections Installed on the leading edge surfaces of the wings, stabilizers, and nacelle inlet lips.
The anti-icing systems use electrical heating elements to heat:
• leading edges of the propeller blades
• No. I, No. 2, and standby pitot static probes
• left and right AOA vanes
• left and right engine intake flanges
• both windshields and the pilot’s side window
Electrically operated windshield wipers remove rain from the windshields.
This system functions automatically when AC power is available. There are two ice Detector Probes (IDP) on the left and right side of the front fuselage
lf either probe detects ice, an ICE DETECTED message is shown on the ED, just below the SAT indication,
If the REF SPEEDS switch set to INCR: the message is shown in white reverse video for 5 seconds, then changes to steady white.
If REF SPEEDS switch set to OFF: the message flashes in yellow.
If both ice detector probes fail, the ICE DETECT FAIL caution light is on. If only one probe fails, is recorded for maintenance.
Airframe deicing can be controlled automatically or manually. Pneumatically actuated rubber deicing boots are bonded to the leading edges of the wings, horizontal and vertical stabilizers, and nacelle inlet lips. Deicing bleed air is ported of each engine an to inflate the boots regardless of the position of BLEED switches. System pressure is regulated to 18± 3psi and shown on the DEICE PRESS indicator, on the copilot’s side panel, A BOOT AIR switch is used to control an isolation valve that interconnects the two systems. It is normally open to ensure uninterrupted operation of either system, in the event of engine failure. The ISO position can be used to check regulated pressure in each system or to isolate a system leak.
Regulated deicer pressure is also used to inflate the forward passenger and aft baggage door seals, and to operate an 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 selector operates deicing, when set to SLOW (3 min cycle,) or FAST(1 min.). The selector is self-homing such that a selection and back to OFF will complete a full cycle. Automatic boot inflation sequence is controlled and monitored by the Timer and Monitor Unit (TMU). The TMU controls the sequence and supplies a dwell period related to the selected rate. Green WlNG, TAIL and nacelle inlet lip boot inflation lights show boot inflation sequence and confirm correct boot inflation pressure.
NOTE -To make sure deice pressure is maintained at 15 psi or greater during decent, holding and approach, it may be necessary to increase NL by advancing the POWER levers.
Integral DDV heaters automatically come on when:
TMU temperature monitor parameter has not failed
AIRFRAME MODE SELECT selector set to OFF, SLOW or FAST
Static Air Temperature (SAT) is less than +5 deg C’.
The DEICE TIMER caution light comes on if there is a failure of the TMU
automatic deice sequence logic
When the AIRFRAME MODE SELECT selector is set to the MANUAL position, the DDV heaters come on permanently.
If a malfunction occurs in the automatic timer or a leak occurs in the system, the boots can he cycled manually with the AIRFRAME MANUAL SELECT selector.. Each set of related boots will inflate as long as the selector stays at the set position on to show full inflation.. A. minimum dwell time of 24 seconds should be observed between each cycle.
If a leak occurs to the pneumatic lines of either deice system, the BOOT AIR switch can be set to ISO, the isolation shutoff valve closes, isolating the failed side from the operating pressure side.
All the boots on tail section are pneumatically cross connected through a restrictor.Thus, if only one side has pneumatic pressure, pressurized air can still pass to the stabilizer boots on the no pressure side. However, the intake boot on the failed side will not inflate.
The DEICE PRESS caution light will come on if the:
• main deice pressure on either side is less than 15 psi
• boot pressure does not reach I5 psi after the DDV opens
• boot pressure stays at I5 psi after the DDV closes
The propeller blade leading edges are protected from ice accumulation by electrically heated elements bonded onto each blade. Electrical power is supplied from the related AC bus.
The propeller deice system is operated by the PROP selector, All six blades on one propeller are heated at the same time during a deicing cycle. To minimize the electrical load on the system, the propellers are alternativly heated. The green PROPS advisory light comes on when the related propeller is heated. The healer is controlled by a Timer Monitor Control Unit (TMCU). The TMCU heater cycle times depend on the Total Air Temperature (TAT).
The TMCU will heat all 6 blades on one propeller when the:
PROP selector is set to ON
TAT is less than or equal to ±50C
Np is above 400 RPM
AC power is available
NOTE - The effectiveness of the propeller deicing system can be improved and propeller vibration reduced by operation of the propellers at 1020 RPM (MAX).
When the PROP selector is held in the spring loaded TEST position, each propeller will be heated for a cycle of 7 seconds, if RPM >400 and AC Available.
The test cannot be started again for 30 seconds.
There are composite fuselage ice protection panels, on both sides of the fuselage, in line with the propeller arc.
PROP DEICE caution light comes on if:
• there is a major failure of the propeller deicing system
• one or more blade heaters are not being heated
• TMCU is not powered
There are two electric heaters, primary and secondary, installed in the intake flange of each engine. The heaters are powered by 115 VAC. Heater operation is confirmed by the HTR segment. (amber), on the ENGINE INTAKE switchlight..
The primary heaters are energized when:
• the engine intake bypass doors are opened
• oil pressure in related engine
• temperature is less than ± 150C
• AC power is available
If the primary heater fails, the secondary heater will he energized. There is no indication of this in the flight compartment. The fault is recorded in the CDS. for maintenance.
If both heaters fail, the related ENS ADPT HEAT 1 or ENS ADPT HEAT 2 caution light comes on.
The engine intake flange is also heated by engine oil. Oil temperature must he kept above 65 deg C (55 on gnd re mod) to make sure there is air inlet ice protection. The engine oil temperature gauge green arc will shift upward to 65 deg C (55 on gnd re mod), if the PROP selector is turned to TEST or ON.
The No. I, No.2 and standby pitot static probes have integral heaters which are activated by the flight Clew to prevent ice build up. The No. I pitot static probe heater is powered from the left AC bus, and the No. 2 probe is powered from the right AC bus. The standby pitot static probe (#3) heater is powered from the 28 V DC Right Essential bus. All three pitot static probes are controlled and monitored by the TMU. The probe heaters are selected ON and OFF by the PITOT STATIC switches.
The related PITOT HEAT STBY, 1, or 2 caution lights come on if
• PITOT STATIC circuit not powered
The left and right AOA vanes are heated whenever AC power is available. They are powered from their respective side AC bus and are directly connected to their power supply through the TMU.
Heater failure is indicated by a PUSHER SYST FAIL caution light, and the related STALL SYST FAIL caution light.
The left and right windshields and the pilot’s side window are heated to supply anti-icing and demisting. When the selector is turned to NORM, each windshield is heated at full power from its related AC bus. When the PLT SIDE WDO/HT toggle switch is set to ON, the forward part of the pilot’s side window is heated from the right AC bus.
Anti-ice controllers control the windshield and side window heaters. An overheat condition shuts off power to the related windshield or side window heater. If either windshield controller fails, the WSHLD CTRL caution light comes on. If either windshield overheats, the WSHLD HOT caution light comes on. If the pilot’s side window overheats, the SIDE WDO HOT caution light comes on.
Each windshield has a wiper which is controlled from a single WIPER selector on the WINDSHIELD panel, with positions PARK, OFF, LOW and HIGH. Turning the selector from LOW or HIGH to OFF stops the wipers at their existing position. When the selector is held at the spring-loaded PARK position, the wipers operate at low speed and automatically stop at the parked position.
There is a guarded ALTERNATE PILOT WIPER switchlight on the pilot’s console. Pushing it in to ON, operates the pilot’s wiper at high speed.
An ice detector spigot is installed on each windshield wiper arm for inspecting possible ice accumulation. Momentary W/S WIPER ICE DETECT pushhuttons, one on each side console, are used to light the related spigot at night.
LANDING GEAR AND BRAKES
The Dash S Series Q400 has retractable dual wheel tricycle landing gear
The landing gear is electrically controlled, hydraulically operated, and mechanically locked. The main gears retract aft into the nacelles and the nose gear retracts forward into the nose section. Gear doors completely enclose the landing gear when it is retracted, and partially enclose the gear when it is down.
Advisory lights show position of gear, gear doors and downlocks. An aural warning sounds if the gear is not extended during certain flap or power configurations.
If the landing gear cannot be extended normally, there is an alternate landing gear extension method. There are also alternate downlock verification lights.
A Proximity Sensor Electronics Unit (PSEU) monitors and controls the operation of the landing gear components. The nosewheels are steerable by a hand tiller and by the rudder pedals.
The main wheels are equipped with anti skid multiple disc carbon brakes. The brakes can be controlled by the brake pedals or the EMERG BRAKE lever.
The main gear retracts aft and has multiple disc carbon brakes with an anti skid system The nose gear retracts forward and has steerable nosewheels The landing gear is operated by the No. 2 hydraulic system and is controlled by the landing gear selector lever on the LANDING GEAR control panel. There is an alternate method for extending the 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 . When the gear is up, all doors enclose the main wheels. When the main gear is down, the hydraulic aft main doors are closed, while the forward main doors stay open. The nose gear has a pair of forward and aft doors, which completely enclosed the nose gear when the gear is up. When the nose gear is down, the hydraulic forward nose doors are closed, while the aft nose doors stay open.
The Proximity Sensor Electronic Unit (PSEU) controls the landing gear hydraulic doors and related advisory lights. It also monitors the Weight-On-Wheels (WOW) proximity sensors. WOW signals prevent gear retraction while on the ground. If the WOW system malfunctions, the WT ON WHEELS caution light comes on. Redundancy is built in to make sure the landing gear operates if there is a sensor failure. If the gear is not down and locked with particular flap or power settings, an audible warning tone sounds.
An integral ground lock mechanism, on the left nose of the aircraft, locks the nose gear. Ground lock pins are supplied for the main gear. The main gear lock-pins may be kept in the MAIN LANDING GEAR STOWAGE of the forward passenger door. With the gear down, the pins are inserted into the main gear stabilizer brace assemblies..There are also landing gear door lock pins for the nose and main hydraulic doors, to keep them open.
Optional tire pressure gauges can be installed on all the tires, with the inflation valves. These show the related tire pressure on a dial, and denote the correct pressure ± 5%.
Landing gear operation is controlled and monitored from the LANDING GEAR control panel. The landing gear is selected UP or DN (down) by moving the landing gear selector lever. A LOCK RELEASE selector lever must be held down to let the gear selector lever move in either direction.
If any of the green LEFT, NOSE or RIGHT advisory lights fail, three green downlock verification lights can be used to check that the gear is locked down. These green verification lights are located in the LANDING GEAR ALTERNATE EXTENSION panel, in the flight compartment floor. The PSEU uses other proximity sensors for these verification lights.
GEAR WARNING TONE
A landing gear warning tone sounds over the flight compartment speakers if the HORN switch is held at TEST. It also sounds if the gear is not down and locked and:
I. • FLAPS lever> 8.5 deg
• either engine torque < 50%
• both PLA < RATING detent
2. • both PLA <FLIGHT IDLE +12 deg
• KIAS <156
• RA < 1053 ft (321 m) if it is valid
3. • one PLA <FLIGHT IDLE ± 12deg
• other PLA < RATING detent
• KIAS < 156
• RA < 1053 ft (321 m) if it is valid
• HORN switch not latched at MUTE
NOTE - No. 3 is the only case in which the landing gear warning tone may be muted.
When the landing gear selector lever is moved to the UP position, No. 2 hydraulic pressure is applied to the retract side of the system The main and nose hydraulic doors open, and the gear retracts. The hydraulic forward nose and aft main doors close after the gear is up.
The advisory light sequence during retraction starts with the LEFT, NOSE and RIGHT red unsafe lights and the amber selector handle light coming on. At the same time, the green LEFT, NOSE and RIGHT lights go off to show the gear is not locked down. The amber DOOR advisory lights come on to show the hydraulic doors are open. When the landing gear is retracted and locked in the up position, the amber selector handle light and red advisory lights go out. Finally, the amber door advisory lights go off to show all the hydraulic doors have closed. No advisory lights should be on if the gear is up correctly. The landing gear, and all hydraulic doors, are held in the up position mechanically. Hydraulic pressure is then removed from the system.
When the landing gear selector lever is moved to the DOWN position, hydraulic pressure is applied to the extend side of the system. The main and nose hydraulic doors open, and the gear extends. The hydraulic forward nose and aft main doors close after the gear is down and locked.
The advisory light sequence during extension starts with the LEFT, NOSE, and RIGHT red unsafe lights and the amber gear selector handle light coming on. The amber door advisory lights then come on to show the hydraulically operated gear doors are open. When the landing gear is fully extended and locked in the down position, the red unsafe lights, and the selector handle light goes out. As each red unsafe light goes out, the related green LEFT, NOSE, and RIGHT advisory lights come on. Finally, the gear door advisory lights go out when the hydraulically operated doors are closed. Continuous hydraulic pressure acts on the doors and gear when down and locked, however, primary downlock is by the overcenter locks.
If a landing gear hydraulic sequencing valve fails, or the PSEU is unable to control the valve, the LOG GEAR INOP caution light comes on.
NOTE - If the LOG GEAR INOP caution light is on, ALTERNATE GEAR EXTENSION procedure is to be followed.
The alternate gear extension system gives a means of extending the landing gear if:
• LDG GEAR INOP caution light is on
• Landing gear does not extend normally
• Loss of No. 2 hydraulic system pressure or fluid
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 prevents the LANDING GEAR selector handle from operating the gear.
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 MAI N L/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 hand pump socket and the NOSE L/G RELEASE handle. Opening the door mechanically operates the alternate gear selector valve, for the main gear. The extension pump handle, located behind the copilot, is inserted into the pump handle socket and operated to complete main gear extension and downlock. 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, and is assisted by airflow to a down and locked position.
NOSEWHEEL STEERING SYSTEM
Directional control on the ground is by the Nosewheel Steering (NWS) system, powered by the No.2 hydraulic system. Steering control is by either the hand tiller or the rudder pedals. The tiller turns the nosewheel up to 70 deg. either side of center for low speed taxi. Steering with the rudder pedals turns the nosewheel up to 8 deg. either side of center for high-speed taxi, takeoff and landing roll. After takeoft the nosewheel automatically centers before retraction.
The tiller, located on the Pilot’s Side panel, is self-centering and operates when the STEERING switch is set to the STEERING position. The nosewheel must be within 70 deg. of center for the steering to operate. An index mark on the hand control shows the relative position of the nosewheel against a fixed STEERING RANGE decal. With the STEERING switch set to STEERING, the nosewheel Steering Control Unit (SCU) is powered, when the nose gear is down and locked with weight-on-wheels.
CAUTION - Do not set the STEERING switch to STEERING if tow bar is connected to nose gear.
CAUTION - Nose steering is not available for approximately 2 seconds after setting the STEERING switch to STEERING.
The nosewheel will revert to a passive shimmy dampened castoring mode if:
• nosewheel angle is more than 70 deg
• SCU detect a steering system fault
• STEERING switch is set to OFF
In the passive mode, the nosewheel will castor up to 120 deg either side of center. Differential braking and/or power may be used for directional control in the passive mode.
The NOSE STEERING caution light comes on if: STEERING switch set at STEERING and
• Steering system fault
• Nosewheel forced more than 70deg
Note A latched fault (Nose Steering CL)will register if tiller is >8 deg and gear down and W of W. ie you tiller befor NW touchdown
• STEERING switch set at OFF
• hydraulic pressure within steering system (Hyd SOV remains Open and may indicate a Hyd lock)
Note The NOSE STEERING caution light does not come on if electrical power is removed from the SCU.
TAXIING IN REVERSE
When taxiing in reverse the STEERING switch must be set to STEERING. However, do not use steering tiller or rudder pedals during reverse taxiing.
CAUTION - Exercise extreme care when taxiing in reverse due to the length of the fuselage.
NOTE - Taxiing in reverse should only be conducted on paved surfaces and in crosswinds less than 10 knots.
Each main wheel is equipped with a multiple disc brake unit powered by the No.1 hydraulic system. An Anti Skid Control Unit (ASCU) modulates the application of brake pressure to each brake unit.
Brake pressure is applied by pushing the pilot’s or copilot’s brake pedals.
• Monitors wheel speed
• Modulates the brake pressure applied to each brake unit to prevent wheel lock-up
• Gives maximum braking at all levels of runway friction.
The ANTI SKID switch on the copilot’s glareshield panel operates the anti skid system when set to the ON position and the wheel speed is more than 10 knots. A start up self test of the anti skid control circuits is made when the switch is moved to the ON, or momentary TEST. The start up self test is prevented with wheel speed more than 17 knots. If the ANTl SKID switch is held at TEST when the aircraft is on the ground, the INBD and OUTBD ANTISKID caution lights come on for six seconds and then go off. If the ANTI SKID switch is held at TEST in the air, with the landing gear extended and locked, the ANTISKID caution lights come on for three seconds. If the ASCU senses a fault in the system, the related ANTISKID caution light comes on.
The PSEU supplies weight-on-wheels and gear up and locked, signals to the ASCU, to make sure that the brakes are off until the aircraft has touched down and the wheels are spinning. The ASCU gives a 5 second delay before brake pressure is applied. This delay is cancelled when wheel speed is more than 35 knots.
NOTE Brake cooling times (Volume 3) must be observed between a landing, or a low energy rejected takeoff and a subsequent takeotf to make sure that sufficient brake energy is available to bring the aircraft to a complete stop if the subsequent takeoff is rejected.
EMERGENCY BRAKE SYSTEM
The emergency/parking brake system is available if the normal brake system fails, or for setting the parking brake. An EMERG BRAKE lever operates the system. The emergency/parking brake system pressure uses the No. 2 hydraulic system, or from the parking brake accumulator. Emergency/parking brake hydraulic pressure is shown on the PK BRK indicator, on the Permanent Systems Area (PSA) of MFD2.
The EMERG BRAKE lever operates against a spring to produce a resistance proportional to the brake pressure applied. There is no differential braking and no anti skid protection. A fully charged accumulator is sufficient for approximately six full applications, in the event No.2 hydraulic system is inoperative.
The parking brake is engaged by pulling the EMERG BRAKE lever to the detent PARK position. The PARKING BRAKE caution light come on when the EMERG BRAKE lever is pulled from the full forward position. The button on the side of the handle must be pushed to release the lever from the PARK detent.
NOTE - Care should be taken when releasing the lever as considerable spring tension will force the lever forward
NOTE - With the parking brake set, application of engine power will cause the takeoff warning horn to sound.
A hand pump, located in the right main wheel well, can be used to increase the park brake system pressure. If external AC power is applied to the AC buses, the standby hydraulic pump and the PTU can be used to increase the system pressure.
There is a fixed oxygen system for the fight crew and observer. Separate portable oxygen systems are supplied for the passengers and attendants. Separate breathing units are supplied for use in low oxygen environments.
Optional First Aid and separate attendant oxygen is available.
The crew fixed oxygen system includes three microphone-equipped masks with an oxygen dilution regulator. Smoke goggles or optional full-face masks are supplied for crew protection in the fight deck. The portable passenger oxygen cylinders are kept in the cabin. Protective Breathing Equipment (PBF.) units are available for the flight crew and attendants. Optional first aid oxygen is kept in the passenger compartment. Optional attendant oxygen is kept in the forward stowage area.
CREW FIXED OXYGEN
1he crew fixed crew oxygen system supplies supplemental oxygen for a descent to 14,000 feet in four minutes and a flight at 14,000 feet for 116 min. The crew masks are supplied from a. single common cylinder, in the right lower nose compartment. A gauge on the cylinder shows the cylinder pressure. A green burst disc, located on the right side exterior of the nose, is ejected out if cylinder over-pressurization occurs
The cylinder pressure is also shown on a lighted oxygen gauge, on the copilot’s side console. If the cylinder is turned off, this indicator pressure shows atmospheric, and will be in the red zone. The cylinder gauge will still show pressure.
The three crew masks are kept in stowage cups, on the bulkhead behind the pilot and copilot’s seats. Each mask plugs into its related oxygen outlet. The observer’s mask, which is kept adjacent to the copilot’s mask, is supplied from the dual outlet on the copilot’s oxygen supply line. The mask supply hose has an in-line pressure indicator, to show if there is sufficient oxygen pressure. The indicator shows green with correct pressure, and red if the oxygen pressure is low. If one mask fails to operate, the observers mask may he used by either crew member.
The oxygen masks have a microphone with an audio connector. If the masks are used with the smoke goggles, they are approved as (Protective Breathing Equipment) PBE. The smoke goggles are kept in pockets in each side console and at the observer’s station.
Each mask has an inflatable harness so that the mask can he put on in less than five seconds. A red inflation button inflates the harness. A rotary three positions knob controls the automatic diluter demand regulator installed on each mask.
• NORM, regulator automatically supplies an air/oxygen mixture
• l00% regulator supplies 100% oxygen regardless of altitude
• EM ER, regulator supplies 100% oxygen at a positive pressure
The EMER position will purge smoke from the smoke goggles.
Smoking is not permitted when oxygen is in use.
Keeping the regulator in the EMER position can deplete the oxygen system
If difficulties in breathing are experienced or the line indicator is red, ensure the supply hose is connected.
There are four Protective Breathing Equipment (PBE) units. In the flight deck, the PBE container is behind the copilot’s seat. In the cabin, the PBEs are located adjacent to the fire extinguishers in the forward and aft storage areas.
The PBE unit is a self-contained, portable-breathing device, stored in a container with a see through door (The PBE is contained in a vacuum-sealed pouch for moisture and contaminates protection. A pull strip on the pouch is used for opening .
The PBE has a hood that seals around the operators neck to keep out gases and fumes. The hood is ready for use if the pull strap is intact and there is no visual loss of vacuum to the pouch. The hood has a clear panel and permits aural communication, such as with the interphone or radio.
The PBE has enough initial oxygen to fill and purge the hood automatically after the hood is put on.
The P81+ produces oxygen when exhaled air passes through a canister. The canister supplies a minimum of I5 minutes of oxygen.
The passenger portable oxygen system has three oxygen cylinders that supply the oxygen in case of depressurization. They are kept in the forward draft bulkhead and floor bins at the aft end of the cabin
The assembly consists oft
Portable oxygen bottle and regulator
Oral/Nasal mask with connecting hose, quick disconnect, indicator, and reservoir bag
Masks tote bag
Inline pressure indicator.
The continuous flow regulator assembly consists oft
• Pressure gauge
• On/off control valve
• Three quick-disconnect outlets for passenger masks
• Overpressure discharge disc
• High-pressure relief valve – 100 psi
. Each cylinder can sustain three passengers for a minimum of 30—minutes.
Smoking is NOT permitted when oxygen is in use.
With the cabin pressurized, this cylinder is unsuitable for first aid use; use first aid oxygen
The portable passenger oxygen supply is not intended for first aid purposes.
Two optional first aid oxygen cylinders, labeled FIRST Aid OXYGEN, are located in the cabin above the lockers between the rearmost seat and bulkheads. Each cylinder is mounted to a bracket assembly, and kept in position by two clamps
The assembly consists of
• Portable oxygen bottle and regulator
• Oral/Nasal mask with connecting hose and inline pressure indicator
• Carrying strap
The regulator assembly consists of
• Pressure gauge
• On/off control valve
• Ill and LOW flow outlets
• Recharging valve
• Overpressure discharge disc
• High-pressure relief valve – 100 psi
Each first aid oxygen mask has a pressure indicator in the supply hose. The indicator shows green when minimum supply pressure is available.
Oxygen is supplied at a constant pressure to the mask when the control valve is turned counterclockwise to the ON position and the mask is connected to either the HI or LOW outlet. The oxygen mask is normally connected to the HI flow outlet, however, it may be connected to the LOW flow outlet to conserve oxygen.
MINIMUM DISPATCH PRESSURE
Minimum dispatch pressure for the oxygen systems is:
Crew fixed oxygen system at 70 F (21~ C)
40 cu ft. cylinder
1300 psi (2 Crew)
1800 psi (3 Crew)
50 cu. ft. cylinder (CR42954!)
1050 psi (2 Crew)
1450 psi (3 Crew)
Passenger portable oxygen system
1600 psi per cylinder
First Aid oxygen cylinder
1650 psi per cylinder
The pneumatic systems use bleed air from the engines or an optional Auxiliary Power Unit (APU).
The pneumatic system supplies air for:
Door seal pressurization system
Engine Oil Cooling ejection system
Bleed air from the Low Pressure (LP) compressor, or the High Pressure (HP) compressor of the left and right engines is ducted to the system. POWER lever position and the Environmental Control System (ECS) Electronic Control Unit (ECU) or pressure switches within the system determine whether the HP port or the LP port will be used.
Bleed air for airframe deicing is also supplied . It is cooled and pressure regulated.
Selecting the BLEED Control switches on the AlR CONDITIONING control panel to I and 2, turns on the related engine bleed air system. Two PACKS switches, labeled OFF/MAN/AUTO, and a single rotary BLEED control selector ( Min., Norm., Max.) control bleed airflow to the ECS. Together they regulate the quantity of air lowing into the system.
Pushing the bleed air (BL AIR) switch light on the APU CONTROL panel turns on the APU bleed air.
The Bleed Air System (BAS) is part of the Environmental Control System (ECS), which supplies conditioned air to the flight deck and cabin. Bleed air flows through ducts from each engine along the wing forward spar, through the dorsal fin then to the air conditioning packs,. The APU can supply bleed air on the ground when the Bleed switches are OFF?
An ECS Electronic Control Unit (ECU), and the selections made on the AIR CONDITIONING panel, control the flow of engine bleed air to the air conditioning packs .
Bleed air is supplied to the deicing system, regardless of the position of the BLEED switches.
At the start of a flight, the ECU selects one digital channel to control the shared components and functions of the bleed air system.
The Environmental Control System (ECU) has a left digital channel with a left backup analog channel and a right digital channel with a right backup analog channel.. The digital and/or analog channels monitor the bleed protective sensors to maintain the correct temperature, pressure, flow rates and valve positions during normal operation. If one of the protective sensors exceed certain values, the ECU automatically shuts the system down.
The ECU responds to selections from the BLEED switches to configure the Bleed Air System for bleed source selection. The ECU and the Bleed Air System have redundant configurations for continued ECS operation with mechanical and electrical component malfunctions.
The ECU uses a flow sensor to measure and balance the bleed air flow through each Nacelle Shut-Off Valve (NSOV)from both engines to the ECS. Bleed sharing can occur only when both engines and both digital channels are operating. When a( the? ) digital channel loses power or fails, or if the PACKS switch is set to MAN, the NSOV fails fully open. There is no analog channel function for bleed air sharing.
Bleed air flow to the ECS is controlled by the BLEED switches and BLEED control selector, on the AIR CONDITIONING panel. After engine start, the bleed control switches are set to BLEED 1 and 2 and the BLEED control selector set to MIN, NORM or MAX, the ECU opens the NSOVs and bleed air is supplied from each engine to the ECS. At high power settings (takeoff climb and cruise) bleed air is supplied from the LP (low pressure) port , at low power settings (descent), bleed air comes from the HP port.
If the BLEED switches are set to OFF, the ECU closes the:
• High pressure shutoff valve (HPSOV)
• Nacelle shutotl valve
The valves will fail safe closed with loss of electrical power, Bleed air to the ECS stops, but still flows to the deicing systems and oil cooler ejector. Bypass circuits to the HPSOV ensure sufficient pressure is available to the deicing system at all times. ???
The BL AIR switchlight, on the APU control panel controls that source. Both BLEED 1 and 2 switches must he set to OFF for it to operate. The BLEED selector has no effect on APU BL AIR flow.
If an overtemperature, overpressure or bleed leak condition is detected, the ECU automatically shuts down the bleed air system, and illuminates the associated #1 or #2 BLEED HOT caution light. The ECU shuts off the associated:
• High-pressure shutoff valve (HPSOV)
• Nacelle shutoff valve
Both lights indicate a bleed leak, and pressurization is at risk.
The BLEED HOT caution light is a latched condition.
Pressure to the deicing system is not affected.
With bleed air shutoff as a result of a BLEED HOT, or a BLEED switch set to OFF, or following an engine failure, bleed flow remaining defaults to a fixed value regardless of MIN/NORM/MAX selection.
Engineering redundancy provides that each digital and analog channel sends a discrete signal to turn on the BLEED HOT caution light.
Two Pratt & Whitney PW I50A turboprop engines power the Dash 8 Q400. Each drive a six bladed. constant speed, fully feathering, reversible Dowty R408 propeller. Each develops 4580 Shaft Horse Power (SHP) under normal takeoff conditions. An automatic uptrim, or optional MTOP rating develops a maximum takeoff power of 5071 SHP.
The engines have a (first stage) axial compressor and a high pressure (second stage) centrifugal compressor, each attached to separate single stage turbines. A two-stage power turbine via a third shaft drives the propeller through a reduction gearbox. The high-pressure compressor drives an accessory gearbox.
POWER and condition levers are used for related engine and propeller control. Full Authority Digital Electronic Controller (FADEC) units control engine power with data from POWER lever position. The POWER levers change engine power in the forward range, and propeller blade angle/power in the idle through reverse (beta) range. The condition levers, through related Propeller Electronic Controller (PEC), set propeller RPM and engine power ratings in the forward thrust range, manual propeller feathering, and fueI on/off control.
Air entering at the engine inlet is ducted to the low-pressure (NL) axial compressor and then to the high pressure (Nfj) centrifical compressor where it undergoes a second stage of compression. It then travels via internal ducts, and is discharged into the combustion chamber where fuel is added and ignited. Gases exiting the combustion section initially impact onto a single stage Nh turbine that extracts energy from the flow, and drives the NH compressor. A gear drive attached to this compressor drives the accessory gearbox mounted on the top section of the engine. Mounted behind the Nh turbine is a single stage NL turbine, that drives the NL compressor. As the combustion gases continue to flow rearward they are directed towards the two-stage power turbine. The power turbines turn as a single unit extracting the majority of gas energy remaining to rotate the reduction gearbox at the front of the engine. Through the reduction gearbox, power is transmitted to the propeller. After leaving the power turbine, the gases flow through the exhaust pipe where they are vented overboard
An accessory gearbox, operates:
• Oil Pressure and Oil Scavenge Pumps
• High Pressure Fuel Pump
• Permanent Magnet Alternator (PMA)
• DC Starter Generator
Each engine nacelle intake incorporates a. bypass door, which provides a means of preventing solids from entering the engine intake. Door opening and closing is controlled by switchlights on the ICE PROTECTION panel.
The doors are selected open when operating in conditions of:
• Heavy Precipitation
• Bird Activity
• Contaminated Runways
• Gravel Runways
The Full Authority Digital Electronic Controller (FADEC) is a dual-channel microprocessor-based controller that controls the engine fuel flow based on various inputs from the aircraft, engine, and popeller control system. The FADEC also controls two bleed valves on the engine for surge avoidance during stable and transient operation.
The engine and controls have sensors providing feedback signals to the FADEC for engine control, cockpit indication, engine health monitoring, and isolation of component failures.
Pressure sensors indicate the status of the engine fuel and oil systems.
Engine information from the FADEC is transmitted to the Engine Display (ED) and provides:
TRQ torque developed within the engine indicated as a percentage of the maxImum.
PROP Propeller speed indicated in RPM.
NH Hp turbine /compressor speed as a percentage of maximrlm speed.
ITT Inter Turbine Temperature shown in degrees Celsius.
NL Lp turbine /compressor speed as a percentage of maximum speed.
FF Fuel Flow shown in hundreds of pounds per hour ( kilograms per hour.)
Dual split analog/digital gauges provide oil temperature in degrees C and pressure gauges in psi
The engine mounted Permanent Magnet Alternator (PMA) provides primary and independent source of electrical power to the FADEC when gas generator speed (NH) is above 20%. The essential power buses provide alternate electrical power to the FADEC for engine starting or PMA malfunctions.
There are two Handling Bleed Valves (HBOV) per engine that provide increased surge margin for engine handling during starting, steady state and transient operation. One is to bleed LP air (steady state operation) and the other to bleed HP air (transient operation).Their operation is automatic.
The Fuel Metering Unit (FMU) (controls the fuel flow supplied to the engine based on demand from the FADEC. The FADEC calculates based on power request and engine sensor inputs such as NH, NL, NP, torque and ambient conditions. The fuel pump is driven by the engine through the accessory gear box and delivers pressurized fuel to the FMU. Excess FMU fuel is returned to the airframe fuel tanks as motive flow for the main /scavenge ejector pumps.
The engine oil system provides lubrication of the engine, gearboxes, and propeller operation. The integral oil tank (approximately 6 US gallons total) in the engine has a sight gauge/dip stick to check oil quantity.
NOTE - Oil quantity must he checked within 15-30 minutes after engine shutdown.
Oil temperature is controlled by the Air Cooled Oil Cooler (ACOC) and the ACOC flap door. The ACOC has an internal thermal bypass valve which controls the oil flow through the cooler The ACOC flap door position is changed automatically, by an electric actuator, to control the amount of air flow through the cooler. Cooling flow augmentation is provided on the ground by use of bleed air.
When the oil pressure drops below 44 psi, a related low oil pressure switch illuminates the ENS OIL PRESS warning light .
Each engine has an
ignition system consisting of one dual channel exciter unit and two ignitor
plugs in the combustion chamber. The system is activated only by the FAD EC during
the start sequence, detected flameout, or surge.
The start system is armed by selecting the SELECT switch, on the ENGINE START panel, to the engine to be started. Pushing the START switchlight causes the starter/generator to rotate the High Pressure (NH) compressor through the accessory gearbox. This develops the necessary airflow and engine RPM, before fuel is introduced. The condition lever is moved 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 conditions and oil temperature.
• When one of the two igniters is turned on , If the ITT does not increase by 200C after 8 seconds of fuel being selected on, the FADEC turns on both igniters. and a records a fault for maintenance record.
• When NH is more than 50%, the ignitors(s) is automatically turned off
• the FADEC controls engine run-up to the requested NH speed, or power
At 50% NH start sequence terminates, the SELECT switch is automatically released to the center position, and the START segment of the START switchlight goes out. Up to fifteen seconds later, the SELECT segment of the START switchlight goes out. The engine continues to accelerate to ground idle, stabilizing above 64.2% NH.
During ground starts only, to make sure that the engine start does not cause over temperature, the FADEC has active ITT limiting to reduce the fuel flow if required (below the standard start schedule). The FADEC will automatically stop fuel flow and ignition if
• the engine does not light within 16 seconds of fuel flow being selected on,
• the ITT limit of 920 deg. C is exceeded:
• NH does not reach 50% in 70 seconds (i.e. hung or slow start).
The start circuitry terminates if NH is above 50%, or the SELECT switch is set to the center position.
ln-flight starts are similar to ground starts except:
• Both ignitors are commanded on
• The auto abort features are disabled
• The FADEC does not actively limit ITT.
FADEC recognizes the difference between an in-flight start and a ground start as 75 KCAS, or with WOW if airspeed is invalid.
A normal shutdown is initiated by moving the Condition Lever to the FUEL OFF position.
A switch in the FMU is energized closed with aircraft electrical power when the PULL FUEL/HYD OFF handle has been pulled.
The PWI 50A engine control is closed loop on power. The FADEC will attempt to eliminate the difference between the requested power (PLA) and the actual power (torque). The authority’ of the power loop is restricted to operational limits.
The power setting logic determines the requested power as a function of engine rating, pilot inputs (such as POWER lever position. [CS bleed selection, etc.), engine failure, and ambient conditions (
The POWER levers allow changes in the power requested from Full Reverse to Rated Power. Ground beta is achieved at PLAs below FLIGHT IDLE. Above FLIGHT IDLE, the power request increases linearly with increasing PLA, until the RATING detent.
Moving the POWER lever in the overtravel region (above the RATING detent) results in an increase in requested power of up to 125 % of the maximum takeoff rating and an increase in the engine software limits. In this region, the propeller control system will automatically set propeller speed to 1020 RPM. This latch can be reset when PLA is removed from the overtravel position and CL is moved.
Rating selection is set with propeller speed selection, when the condition lever is moved to the detent positions:
Condition Lever Posn Rated Pwr
Normal Takeoff NTOP)
900 RPM MaximumClimb(MCL)
850 RPM MaximumCruise(MCR)
START & FEATHER Normal Takeoff
FUEL OFF None
For all condition lever positions, the Rated Power is achieved when the PL s in the RATING detent.
Alternate combinations of propeller speed and engine power rating can be selected by using the MTOP (optional), MCL, and MCR rating discretes pushbuttons, on the ENGINE CONTROL panel. These discretes, override the rating normally selection to the FADEC as a function of condition lever position, under certain conditions.
If the condition lever is in the 1020 RPM position and the optional Maximum Take-Off Power (MTOP) pushbutton is pushed, the rating is selected by the FADEC. The MTOP rating is defined as the maximum available power certified for take—off operation. (NTOP is normallv selected by the FADFC when the condition lever is in the 1020 RPM position.)
When the condition levers are in the 900 RPM position and the MCR pushbutton is pushed, the MCL rating normally associated with this propeller speed is overridden and latched. Subsequent movement of the condition lever releases the latch. Alternatively, the MCL rating can be recovered by pushing the MCL pushbutton.
Selection of MCL, with the condition levers at the 850 RPM position, is also possible using the MCL pushbutton.
The FADEC’s rating selected information is provided to the crew in the top corners of the Engine Display (ED)
This is followed on the next lower line by the Torque bug numeric value (power available)
Below this bleed selection information is provides as follows
Blank – Bleed off
Bleed (white) – Bleed on and Min
Bleed (yellow) - Bleed on and > Min.
When in NTOP, reduced take off power can be selected by pressing the DEC pushbutton, on the ENGINE CONTROL panel, RDC TOP section. REC TOP will replace the NTOP indication, and the new rated power available will be shown on the second line top corner of the ED. Each press of the button reduces rated power by 2% to a limit of 10%. The RESET pushbutton resets NTOP at any time.. If an uptrim is commanded with reduced NTOP, the reduced power will apply to the MTOP requested power, since MTOP will be still only increase NTOP by +/-10%.
Powerplant operation is managed by POWER and condition levers (CL). FADEC and PEC electronically control the engine and propeller, based on the Power Lever Angle (PEA), condition lever position and operating mode.
The power lever (marked I and 2) is used to initiate power demands via electrical signals the FADEC to meter fuel to the engine. The power lever system also initiates control signals to the Propeller Electronic Control (PEC) to control propeller blade angles in the beta range. The FADEC modulates power directly proportional to Power lever movement forward between FLIGHT IDLE and the RATING detent. A flight idle gate prevents unintentional movement of the power levers below FLIGHT IDLE. This is overridden by raising gate release triggers (below the handgrips)..
NOTE - A Beta warning horn will sound if the gate release trigger is raised in flight.
Power lever movement aft of Flight Idle moves the propeller blades into reverse until the POWER levels reach MAX REV. Between DISC and MAX REV, fuel flow and power output is increased.
Two condition levers (CL.) marked I and 2, are used to set:
• Maximum Propeller RPM (MAX/i 020)
• Climb Propeller RPM (900)
• Minimum Propeller RPM (MIN/850)
• Fuel On & Propeller Feather (START & FEATHER)
Engine Shutdown (FUEL OFF)
• Engine Rating
Both the PEC the FADEC receives electrical signals from the CL.
In the constant speed range. with the following default engine rating. With the CL in the MAX/1020 position, the engine rating is NTOP. In the 900 position, the engine rating is MCL, in the MIN/850 position, the engine rating is MCR.
On the ground, with the CL at MIN or above, and the POWER levers at FLIGHT IDLE, propeller is 660 RPM. This is prop underspeed governing. With the CL at the START & FEATHER position, the propeller is feathered, and ground idle operation (220 prop RPM) of the engine is scheduled. Moving the CL to FUEL OFF causes the FADEC to stop fuel flow to the engine.
Lift gates prevent unintentional movement of the CL back to START & FEATHER, and from START & FEATHER to FUEL OFF.
The FADEC discriminates between single and dual engine ECS bled as follows:
Flow is reduced when an engine fails on takeoff (MTOP via uptrim) or a bleed selected off.
Additionally if one engine is operating at MTOP rating, during UPTRIM, setting the related BLEED switch on will change the engine rating to MCP.
An Automatic Takeoff Power Control System (ATPCS) automatically “uptrims” the power of the opposite engine. The failed engine’s PEC/AF sends an Uptrim to the live engine ‘s FADEC which then by changs its engine rating from NTOP to MTOP.
Autofeather armed when both PLAs are “high” and engine torque is at least 50%.
Uptrim is triggered (regardless of AUTOFEATHER selection) if:
• torque of the failed engine falls below 25%, OR
• Np (as indicated by the torque sensor) falls below- 816 RPM, AND
• PLA is set high, AND
• MTOP (optional) is not set.
Either ot the first two condition must be confirmed by both torque sensor signals. The low speed condition accommodates the failure case of a propeller autocoarsening or inadvertently feathering, causing loss of thrust but not low torque. Uptrim is also directly’ signalled if an autofeather occurs. Uptrim signals are sent to the FADEC of the operating engine to increase its power by approximately 10%.
An Uptrim condition is indicated by:
• UPTRIM indication on the ED
• a change in the engine rating from NTOP to MTOP
• a change of the torque bug to indicate an approximate 10% power availability increase
If an uptrim is commanded with reduced NTOP, the reduced power will apply to the MTOP requested power, since MTOP will be still only increase NTOP by 10%.
The Npt Underspeed Governing is used to limit the propeller speed to a minimum of 660 RPM in the air and on the ground. The control system then closes loop on propeller speed and determines the gas generator speed to set the required Npt.
The Npt Overspeed Control Limit in the FADFC prevents the power turbine speed from exceeding 110% (1120 RPM). The FADED signals the FMU to reduce fuel flow.
There is independent overspeed (0/S) protection circuitry (dual channel) built into the FADEC which controls the Fuel Shutoff Solenoid.. A shutoff command is issued when NH exceeds 108%. The 0/S Protection circuitry is exercised on normal shutdowns by the FADEC..
The FADEC prevents engine torque from exceeding limits.
Generally torque is limited to 50% in reverse, 106% in the forward power range, and 125% in the overtravel range.
During such events as caused by a unsheduled feathering of the propeller at high power, The FADEC control via Nh should prevent torque of more than of 135%.
The FADEC annunciates faults depending on their effect on the system. A detected fault is classified into one of three (3) categories:
ENG FADEC FAIL warning light.
The FADEC will:
• stabilize the engine at FLIGHT IDLE or Ground Idle (DISC) depending on the airspeed/W0W OR
• shutdown the engine (commanded by the control system).
No pilot input is required (re the PLA or the CLA).
ENG FADEC caution light.
FADEC operation is limired:
• Asymmetric power levers may be required to obtain symmetric power/thrust.
• Rapid power levers movement may cause engine surge.
This is defined as a fault which is automatically accommodated and is not classified as above. The FADE/C transmits the advisory fault codes to the Engine Monitoring Unit (EMU) and to the ED.
The constant speed six-blade propeller, is fully feathering, reversible, and counter-weighted. Electrical heater elements are bonded to the inner leading edges. The propeller blade is made from carbon and glass fibre-reinforced plastic, with a lightweight foam filler The propeller is driven through, a reduction gearbox, which also drives the:
• Pitch Control Unit (PCU)
• PCU Oil Pump and Propeller Overspeed Governor
• Main Hydraulic Pump
• 115 VAC’ Variable Frequency Generator
The propeller control system consists of the:
• Propeller Electronic Control (PEC) Unit
• Pitch Control Unit (PCU)
• High Pressure PCU Oil Pump and Propeller 0/S Governor Units
• Alternate Feathering Pump
• Propeller System Sensors
The propeller control system adjusts blade angle to control propeller RPM in flight and in reverse. It provides smooth thrust transitions with PLA changes and limits the minimum blade angle in flight. It enables manual and automatic propeller feathering and minimizes cabin noise in synchrophase mode.
The PEC unit is a dual channel microprocessor/controller, located in the engine nacelle. The PEC uses inputs from the aircraft, propeller controI sensors, and the engine control system to control propeller pitch and speed
Each unit performs a number of safety functions isolated from basic control functions. These are:
UPTRIM ( to FADEC of the operating engine.
PITCH CONTROL UNIT (PCU)
The PCU is a hydromechanical device that interfaces with the propeller. Commanded electrically by the PEC, the PCU meters high pressure engine oil to the PCU which controls the flow into the fine or coarse pitch chambers of the propeller pitch change cylinder, so that the blades move in the necessary direction. If the electronic control fails the PCU defaults to the fine pitch position wherein the overspeed governor then controls the propeller and the associated PEC caution light. A complete loss of power will result in a course command and an uncommanded feather scenario.
The High Pressure PCU Pump/Propeller Overspeed Governor (OSG) Unit provides the PCU with high pressure oil from the engine gearbox. The High Pressure pump is driven from the reduction gearbox. The Propeller Overspeed Governor Unit is an independent mechanical system which limits propeller overspeed. The OSG is a flyweight design, driven directly from the drive gear of the pump.
The Alternate Feathering Pump provides an independent means of feathering the propeller. It supplies a secondary source of pressurized oil for feathering the propeller. The pump is 28VDC electrically driven. It is energized by the opposite side secondary bus power through a 30 second relay.
The propeller system is fitted with a Magnetic Pickup Unit (MPU) to provide a signal to the PEC. The PEC uses this signal for propeller speed governing, synchrophasing, and when sent to Active Noise and Vibration Control (ANVC) system for propeller balance information.
Modulating high pressure oil on either side of the pitch change piston controls pitch. Each blade is fitted with a counterweight, around the blade root outer sleeve so that the blade’s natural twisting moment is towards coarse pitch. If modulating oil pressure is lost in flight, the propeller will therefore assume a safe coarse pitch (feather), which causes minimal windmilling drag.
If this oil pressure is lost in reverse, the pitch will change towards the maximum reverse blade angle.
The PCU provides for governed constant speed operation through a propeller governor controlled by the condition levers, at high power. The POWER levers control minimum blade angle in the beta range, to give smooth thrust response. The manual feather mode is controlled by the condition levers or by the autofeather/alternate feather system.
In flight constant speeding operation, the PEC meters sufficient HP oil into propeller fine pitch. This balances the net coarse-seeking moment applied to the blades by the blade counterweights, to achieve the selected propeller RPM. If the HP supply is lost, the blades will “autocoarsen”.. Additionally the PEC responds to acceleration as well as RPM error. (During a slam acceleration, when propeller RPM is undespeeding yet accelerating rapidly to the selected speed, the PLC will direct the servo valve coarse, so that the propeller can absorb more power.)
Constant speeding mode is entered when the propeller speed reaches 850, 900 or 1020 RPM, according to which the condition lever sets RPM. HP oil for constant speeding passes through the OSG prior to PCU control. Should the PCU stick at a fine pitch selection, propeller rpm will increase until approximately 104% (1070 RPM), when the OSC will start to isolate the propeller control system from HP oil,and a stable governing condition at 104% (1060 RPM) will be quickly achieved. Safe overspeed governing is therefore provided, regardless of PEC failure.
The OSG can he tested on the ground by holding the PROP O’SPEED GOVERNOR test switch, on the Pilot’s Side Panel, at TEST.
The Beta mode range is from a PLA just above FLIGHT IDLE (flight beta), to MAX REV. When the propeller is in the Beta mode, blade angle is set by the POWER levers.
During on-ground ‘beta control’ (POWER lever below FLIGHT IDLE), the power lever (PLA) additionally directs the PEC/PCU to meter HP oil to achieve the desired blade angle. The PLA/FADEC/PEC/PCU system operates in closed loop blade angle control. The rate of change of blade pitch is limited to prevent overtorques , over and underspeeding .
During approach, when airspeed and power is low, the propeller will enter flight beta control; entry into the ground beta range is prevented. When the propeller speed increases, the PEC automatically re-enters constant speeding mode.
A flight fine ‘stop’ of 16.5 deg blade angle., is programmed into the PEC and is operative while the POWER lever is at or above FLIGHT IDLE. (ie - in normal flight operation, pitch does not fall below 16.5deg).
If the pitch control fails. The OSG’s hydraulic cut-off’ of HP oil constitutes a flight fine stop interlock that occurs at 16 deg.. This “Flight Fine Stop” keeps a minimum pitch consistent with a positive counterweight effort towards coarse pitch, thus ensuring the effectiveness of the OSG throughout the in-flight range.
To enable blade angles less than 16.5 deg. on the ground, the POWER lever must he brought back below FLIGHT IDLE. A gate on the POWER lever quadrant prevents unintentional movement below FLIGHT IDLE. Blade angles lower than 16 deg. causes the PROPELLER GROUND RANGE lights to come on.
PLA movement into ground beta locks out the OSG. This lock out (valve) position is checked in the OSG test.
(Note OSG Test is a Maintenance requirement.)
With POWER lever is in the beta range, propeller speed NP is controlled by the FADEC/ engine fuel system at a minimum 660 rpm (Np underspeed governing). As Prop Speed is engine driven rather than airspeed driven, the FADEC controls engine speed, and provides protection from the high torque that would result from an inadvertent propeller feather.
Inadvertent Ground Beta PLA in Flight
A software invoked Ground Beta Lockout system prevents blade angles below flight fine, if the POWER levers are moved below FLIGHT IDLE in flight. If either POWER lever is inadvertently moved below the FLIGHT IDLE gate in flight, it should be moved back above the gate.
NOTE - A Beta warning tone will sound if the gate is raised in flight.
In this mode the system operates in closed loop, maintaining the propeller speed between 660 and 1020 RPM. The PLA selects a blade angle and the FADEC schedules fuel to maximum limit (MAX REV) of 1500 Shaft Horse Power (SHP).
This consists of propeller overspeed governing and a “fuel topping” section.
As Discussed the OSG controls blade angles hydraulically, by stopping the HP oil supply, when the RPM exceeds approximately 1060.
When RPM decreases below the overspeed point, the overspeed governor restores normal propeller control
The FADEC Np overspeed circuitry signals the FMU to reduce the fuel supply to the engine, if an overspeed of approximately 1120 RPM is reached. When Np drops below the overspeed point, the FMU to restores normal fuel flow.
This OSG is locked out in reverse, and the FADEC is the primary means of protection from overspeed in the reverse range.
When the speeds of both propellers are within a predetermined difference of each other in flight, the PEG enters a synchrophasing mode to reduce propeller noise, The synchrophasing system does not operate at 1020 RPM. The phase angle is calculated by timing the differences between the master and slave propeller signals over a complete propeller revolution. The phase demand is determined by the condition lever position.
Propeller feathering systems provide
• Alternate feathering and unfeathering
• Manual feathering.
The autofèather system automatically feathers the propeller if and engine fails during takeoff. The system increases the power (uptrim) of the operating engine. Autofeather is selected by pushing the AUTOFEATHER switchlight, on the PROPELLER CONTROL panel. This causes the SELECT light to come on, and A/F SELECT appears on the ED. NP ARM appears on the ED when both engine torques are more than 50% and both POWER levers are more than 60 deg. PLA. If one propeller is autofeathered, the auto feather function of the other propeller is automatically disarmed.
NOTE - If A/P ARM advisory does not appear on ED, take off must he rejected.
Autofeather is activated, from the armed state, when the torque signal of one engine falls below 25%. After a three second delay, the propeller of the failed engine will feather. AN alternate feathering pump is also activated for approximately’ 30 seconds and is supplied oil from an auxillary oil reservoir built into the propeller Reduction Gear Box (RGB), in the event there is no engine oil pressure. The related feathering pump advisory light, in the FTHR switchlight, comes on when the feathering pump electrical contactor is closed.
The autofeather system can he disarmed by:
• Pushing the AUTOPEATHER SELECT switchlight off
• Moving one or both POWER levers to FLIGHT IDLE
• autofeathering of a propeller.
CAUTION - Propeller may unfeather if AUTOFEATHER switchlight is selected off before condition lever is selected to FUEL OFF.
If Uptrim occurs independent of autofeather, it may only be disarmed by moving both POWER levers below the 60 deg PLA (2/3 fwd travel) position.
An Np underspeed cancel signal to the feathering engine prevents the FADEC from increasing Nh (fuel) in an attempt to maintain propeller RPM as the feathering propeller RPM decreases below 660. (Note Flight Idle feather is approx 220 RPM)
The AUTOFEATHER switchlight is also used for testing the system.
Operational back-up/alternate feathering is available by pushing the guarded #1 or #2 ALT FTHR switchlight, on the PROPELLER CONTROL panel, when the condition lever is at START/FEATHER or FUEL OFF position.
Selection of the switchlight;
provides direct access to the DC powered Feathering Pump mounted on the ROB,
operates a back-up feather valve in the PCU and
delivers HP oil to the propeller so it may be feathered regardless of failures in the normal control system.
This feather pump:
• gives a back-up feather function if the primary feathering system is inoperative.
• enables the propeller to be feathered when the gearbox rpm or oil level is too low to maintain a supply to the HP pump.
• enables the propeller to be unfeathered on the ground for maintenance.
Propeller manual feathering is used during engine shutdown by moving the related condition Iever to the START & FEATHER or FUEL OFF position .
The PEC continuously monitors itself and connected equipment for failures. Fault codes are sent to the FADEC. Detected faults can be critical, cautionary, or advisory.
Critical failures involve malfunctions which results in a total or partial loss of’ propeller control.
If there is no electrical power, loss of output from both PEC control chanels or failure of both servo valves, results in a coarse pitch selection (feather). The FADEC also stops the fuel supply to the engine.
Loss of HP oil during flight, the blade counterweight effort will drive the propeller coarse:
on the ground, the final pitch will be indeterminate due to the variable nature of the blade forces at low blade angles.
Cautionary faults provide PEC caution light. With either caution light on, beta control below FLIGHT IDLE is not available.
Faults that have implications for subsequent aircraft dispatch cause a POWERPLANT message to appear on the ED.
Advisory faults are automatically accommodated by the PEC, have no associated operational impact: and will normally be annunciated on request for maintenance only.
AUPC is an independent control circuit protecting against failures that drive the propeller(s) in the coarse direction (feather). Activation is indicated by a PEC caution light(s)
AUPC is armed by PLA at;
FLIGHT IDLE or above,
Condition lever above START FEATHER, and
Torque above 50%, providing both Autofeather and Alternate Feather are not active.
It is triggered if;
Np drops below 816 rpm AND
torque remains above 50%.
When ALPC triggers, an drive fine signal is sent to the PCU, causing the pitch change mechanism to be driven to the pitch line stop. At higher airspeeds, this will result in OSG operation (1070 rpm). The circuit is latched.
Autofeather, has higher priority than the AUPC. i.e. AUPC can not overide the autofeather function.
The AUPC circuitry is confirmed by the A/F TEST.
The aircraft is pressurized by engine bleed air supplied to and distributed by the airconditioning system. Pressure is maintained and controlled by the cabin pressure control system that governs the rate of outflow from the pressurized areas of the aircraft. An aft outflow valve primarily controls the outflow of air, and is assisted by two safety valves.
The outflow valves are controlled from the CABIN ALTITUDE control panel, on the overhead panel. There are independent indicators to monitor the system. The aft outflow valve and aft safety valve are located on the aft pressure dome. A forward safety outflow valve is located on the forward pressure bulkhead.
If cabin altitude is too high, a Cab Press warning light flashes.
Except for the pressure relief functions, the electrically operated aft outflow valve controls pressurization. It is used for automatic and manual control, and can also be used to dump pressurization. The forward safety valve is for emergency operation and for smoke removal. The safety valves have positive and negative pressure relief.
The pressurization system can be controlled in:
A self-test is accomplished when electrical power is first supplied to the system. A FAULT light, on the CABIN ALTITUDE control panel indicates a failure.
With the system in AUTO mode, a preprogrammed Cabin Pressure Controller (CPC) does all pressure scheduling from takeoff’ to landing with minimal input. The computer receives Inputs from the crew and various aircraft systems, and modulates the outflow valve to keep a fixed schedule of cabin altitude versus airplane altitude for complete regulation of cabin pressure during flight.
When the aircraft is on the ground with PLAs at less than 60 degrees, the outflow valve is in the open position to prevent aircraft pressurization. The aft safety valve also opens with at least one engine, or APU operating.
With PLAs greater than 60 degrees (2/3 travel), the controller signals the aft outflow valve to modulate, for the following
• Flight abort after Takeoff
The aft outflow valve moves from the fully open position and starts to modulate. After takeoff (as sensed by the PSEU), the valve set to accommodate the cabin fixed schedule. With bleed air off, the outflow valve and aft safety valve close.
During takeoff with Bleed Air on (using Supplement 21) the cabin is pressurized to 400 ft below the takeoff altitude at a rate of 300 fl/mm. This is to avoid a cabin pressure “bump “.
The CPC retains takeoff mode for approximately 10 minutes after lift off. This avoids the requirement to reselect the landing altitude in case of an aborted flight and return to the departure airport. The pre—pressurization remains in effect as long as:
• The scheduled cabin altitude is higher than the theoretical cabin altitude, or
• The aircraft altitude is less than the takeoff altitude plus 5,000 ft (valid only for takeoff altitude over 8,000 fit
Once one of the above conditions is met, the CPC begins flight scheduling.
When the takeoff sequence is over, the cabin pressurization is in accordance with the preprogrammed pressurization schedule.
During descent, the cabin rate of change is achieved automatically. In the case of a high rate of descent, a descent increase sequence is initiated.
If the set field altitude is higher than actual field altitude, the aircraft will land unpressurized. If the aircraft lands pressurized, the cabin altitude will go hack to field altitude at a programmed rate for one minute, then the outflow valve, and the aft safety valve are fully opened and cabin pressure is bled to ambient.
In the event automatic pressurization mode fails or the FAULT light comes on, pressurization can be manually controlled through the aft outflow valve. The AUTO/MAN/DUMP switch is set to MAN and then the MAN DIFF toggle switch is used. When held to the DECR position, the outflow valve opens and decreases the cabin differential pressure. When moved and held to INCR, the aft outflow valve closes and the cabin differential pressure increases.
NOTE - When operating in manual mode, the cabin altitude, cabin differential pressure, and cabin rate of change indicators should be monitored carefully.
The AUTO/MAN/DUMP switch set to DUMP fully opens the outflow valve,, preventing the aircraft from pressurizing. DUMP mode may also be used for maximum aft smoke evacuation
Pressurization can be controlled through the forward safety valve when the outflow valve becomes unserviceable or undesirable. Pressure can he regulated by turning the FWD OUTFLOW knob, as necessary, to adjust the amount of pressure required. When the control knob is turned clockwise the forward safety valve opens. Pressurization can also be reduced rapidly by turning the forward safety valve selector, on the copilot’s side console, to OPEN. This opens the forward safety valve fully.
NOTE - When operating in emergency mode, the cabin altitude, cabin differential pressure, and cabin rate of chance indicators should be monitored carefully.
Electrical power is not required in this case. Operation is through suction only, created by aircraft slipstream.
The forward safety valve may be used with the automatic system operating during flight to evacuate smoke from the flight compartment. With the AUTO/MAN/DUMP switch set to AUTO, the system will begin to close the aft outflow valve.
During unpressurized flight the cabin and flight compartment can be ventilated with outside ram air. With Bleed Air and Recirc Fan off ram air enters via the dorsal fin NACA vent, through a check valve and into the air conditioning ducting downstream of the packs. With the AUTO/MAN/DUMP switch is set to the MAN position, and the MAN DIFF switch held at INCR (50 seconds), ram air ventilates the cabin and flight compartment and then exhausts through the forward outflow valve,
The maximum differential pressure permitted by the CPC is 5.46 psi, giving a cabin altitude of 8 000 ft at 25,000 ft. alrcraft altitude., Both the forward and aft safety valves release at 5.8 ± 0.15 psi. . Both safety valves also have a negative pressure relief capability that will operate at - 0.5-psi differential to prevent external atmospheric pressure from being more than internal cabin pressure.
The CABIN PRESS vvarning light will come on if cabin altitude is more than 9 800 feet.
Avionics pitot static