AIR CONDITIONING GENERAL
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:
-BLEED selection-MIN/NORM/MAX
-Environmental conditions
-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 detected
Gen. overheat
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.
Capt 7
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:
- Battery
- 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.
NOTE
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;
Batteries
Engine Generator
APU Generator
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)
NOTE
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
-AC External
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:
-Deicing/Anti-icing heaters
-Standby hydraulic pump
-Galley loads
-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.
There are;
- 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:
Engines
APU
Baggage compartments
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:
- Flight
- Ground
- Taxi
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:
outside the TO range,
weight on the nose wheel and
torque above 50%,
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.