D 8 300 manual
basic text write down with hyperlinking

Air conditioning
Brakes normal
-Brakes emergency park prake
Fire protection
Flight controls
Flight instruments
-----Electro mechanical
Ice and Rain
Landing gear
Lighting interior
-Lighting exterior
--Lighting  emergency
Nose wheel steering

The two air conditioning packs condition the bleed air to the proper temperature and humidity. The air is then distributed for environmental control of the flight compartment and cabin.

An optional ground air conditioning connection is available to supply conditioned air from a ground source.

The system uses bleed air from the engines, or from the optional Auxiliary Power Unit (APU), to power two air conditioning packs, that supply the air recirculation system. Two recirculation fans help to boost the supply of conditioned air, and a gasper system distributes air to individual outlets in the cabin, flight compartment and the lavatory. Flight compartment and cabin temperatures are controlled by automatic or manual selections.

Each air conditioning pack has an Air Cycle Machine (ACM), with primary and secondary heat exchangers. Both packs are located in the aft fuselage equipment bay, and cool the hot bleed air from the engines or optional APU.

Bleed air from the engines or optional APU is conditioned by two air conditioning packs, located aft of the pressure dome. The bleed air entering the packs is controlled by the BLEED 1 and 2 switches and the BLEED flow control knob, on the AIR CONDITIONING panel

Both packs have an Air Cycle Machine (ACM), primary and secondary heat exchangers, an air mixer and condenser unit, a ram air exhaust duct, and related piping and control valves. Each pack is supplied with hot bleed air ducted in along the dorsal fin.

Each pack operates independently of the other. Both packs operate on the same principle. Bleed air is routed through the first stage of the heat exchanger where it is cooled before entering the compressor section of the ACM. It then returns to the heat exchanger for two more stages of cooling, before entering the turbine section of the ACM, where the temperature is considerably reduced. This cool/cold air enters a mixing box where it is blended with cabin air forced in by the air recirculation fan and hot bleed air to achieve the desired temperature.

The heat exchangers are cooled by ram air, ducted into the rear fuselage by dorsal fin ram inlets (supplemented by a suck-in door when the airplane is on the ground), and drawn into a ram air duct, in the air conditioning bay, by an ACM driven fan. The duct directs the air through the primary and secondary heat exchangers before discharging it overboard.

Any condensation in the air is extracted by a condenser, and is then sprayed on to the inlet of the heat exchanger to improve its cooling efficiency.

With the PACKS switches in the OFF position, both Pressure Regulator and Shutoff Valves will be closed. There will be no bleed air supply to the packs.

With the PACKS switches in the AUTO position, rotary selectors are used to vary temperature settings. The sensed and selected temperature for each compal1ment is processed in its related automatic temperature controller, which then signals the pack temperature control valve. The pack temperature control valve then establishes the mixture of conditioned and bypassed air for that compartment.

With the PACKS switches in the MAN position, spring loaded toggle switches are used to maintain the desired temperature These switches are labeled COOL and WARM. Flight crew input is required to maintain the desired temperature. The temperature gauge on the AIR CONDITIONING panel shows one of three temperatures

When operating the flight compartment or cabin in manual mode, care must be taken to avoid excessive COOL or WARM selections, that can cause duct over temperature or duct icing to occur.

Varying the blending ratios of cold air to hot bypassed air controls the required temperature. Each pack has related temperature control valves, one in the ACM supply duct, and one in the ACM bypass duct. The valves regulate the blending ratios in response to commands from the related flight compartment or cabin temperature controller.

Cabin temperature can also be controlled from the flight attendant's panel, if the CABIN selector is in the F/A detent position

The air recirculation system supplies air to the mlxing box where it is blended with conditioned and/or hot bleed air to reduce the bleed air requirements for air conditioning. Air is drawn, from under the flight compartment and forward cabin floor, into a duct by two electrical recirculation fans which forces it into the related mixing box. A check valve in the air recirculation duct prevents back-flow when the fan is off.

The recirculation fans are controlled by related RECIRC F/C and RECIRC CABIN switches, on the AIR CONDITIONING panel. Electrical power is from the 28 VDC secondary buses.

If a recirculation fan has overcurrent, over-temperature or undervoltage conditions, it will automatically stop. Automatic restart usually occurs if the condition is cleared. If the fan goes off and stays off it may have actually failed. or it may be latched off. To release the latch, the related RECIRC switch must be set to OFF for a minimum of 20 seconds.

During normal operation, the forward pack supplies air to the flight compartment, and the aft pack supplies air to the cabin. If either pack fails, air to the flight compartment and cabin will continue from the operating pack. Flight compartment air is routed through a baffle duct into underfloor distribution ducts, supplying side vents, gaspers and floor outlets. Cabin air is routed through a baffle duct into underfloor distribution ducts, supplying outlets in the cabin ceiling and floor.

Conditioned air is supplied to the flight compartment to maintain a comfortable environment for the flight crew, side window demisting and aircraft pressurization.

The supply of conditioned air for the flight compartment, enters the cabin at the centre of the rear pressure dome and is routed to the flight compartment through a duct under the baggage compartment and cabin floor.

At the flight compartment bulkhead, the duct splits supplying the airflow into two individual but identical distribution systems, one for the left side and the other for the right side, lower level and upper level outlets. The upper level outlets are demist nozzles for the crew's side windows. The lower level outlets include a foot warming tube (near the rudder pedals), a fixed grille near knee height and a large torso gasper.

Related levers, on the flight compartment sidewalls, control airflow to the side windows and low level vents. The levers are labeled SIDE WINDOW DE-MIST and LOW LEVEL VENTS, and slide from CLOSED to OPEN positions.

The supply of conditioned air for the cabin enters the cabin at the centre of the rear pressure dome and is routed under the baggage compat1ment, where it splits into an upper and lower duct for each side of the fuselage.

The upper duct supplies air to the cabin dado panels, while the lower duct supplies air, through sidewall risers, to the upper air gaspers that are located outboard of the Passenger Service Units (PSU).

Gasper outlets are located at each passenger seat, the flight attendant's station, and the lavatory. The airflow from the outlets can be individually controlled.

When the air conditioning is in manual mode ONLY , and a high degree of cooling is selected, a condition can occur where the air entering the cabin is cooled below freezing. This can result in ice formation that can cause blockage and/or damage to the system.

The Cold Limit Clamp, applies a warmer command to the temperature control system to return the cabin air supply to above freezing. The Cold Limit Clamp works automatically, and requires no flight crew input.

If one pack fails, the remaining pack can supply the entire airplane through the bafi1e box.

If an over temperature condition is sensed in the ducting, the related CABIN DUCT HOT or FLT COMP DUCT HOT caution light will come on. At the same time the related temperature control valve will automatically go to the full cold position. If the condition clears, the light will go out and normal system operation will be restored.

If an ACM compressor discharge temperature exceeds a preset value, the related CAPBINPACK HOT or FLT COMPT PACK HOT caution light comes on. At the same time a signal will automatically be sent to shut off the hot bleed air supply. If the condition clears, the light will go out and normal system operation will be restored.

If both air conditioning packs fail, outside ram air can be used for cabin and flight compartment ventilation. Opening the forward outflow valve using the selector on the copilot's side console will allow maximum ventilation. The airflow enters through the NACA vent on the dorsal fin and exhausts through the forward outflow valve. Refer to Chapter 18 for more information.

The optional Auxiliary Power Unit (APU), is located in the rear unpressurized equipment bay of the aft fuselage. It consists of a gas turbine engine driving a DC starter-generator. The APU supplies bleed air to the air conditioning pack and deicing system, and 28 VDC to the electrical system. The APU cannot be operated in flight.

An Electronic Sequence Unit (ESU) provides automatic control of start sequencing, normal operation, and malfunction monitoring of the APU. The APU starter/generator can be powered from either the aircraft main battery or external power. Intake air is drawn through a screened inlet duct on the right rear of the fuselage. Exhaust gases are vented through an outlet on the right rear fuselage, aft of the intake. The APU is protected by its own automatic fire detection/extinguishing system that continuously monitors the APU and its compartment when electrical power is supplied. The APU CONTROL panel is mounted on the overhead console.

APU fuel is supplied from the left wing collector bay through an APU shutoff valve. A gravity-fed, APU-driven fuel pump maintains positive fuel pressure supply.

The APU shutoff valve opens when the APU PWR is powered. Green and white FUEL VALVE lights, on the APU Fire Protection Panel (FPP). show when the valve is open or closed. Fuel is automatically scheduled for starting. acceleration. and speed regulation by the ESU.

The APU shutoff valve will close if:
• APU PWR switch light is pushed off
• FIRE TEST pushbutton is pushed
• APU automatically shuts down
• Fire is detected in the tail cone
• Aircraft is off the ground.

The APU starter-generator system consists of a gearbox mounted 28 VDC, 300 amp starter-generator and a Generator Control Unit (GCU). When the APU is operating. the RUN segment comes on to show that the generator mode is available to supply 28 VDC. The starter generator. in generator mode. supplies 28 VDC to the right main feeder bus. The GCU closes necessary bus ties and connects all DC buses to the APU. The WRN segment of the GEN switch light indicates the generator is not supplying power.

After the aircraft main engine DC starter-generators are on line. The APU generator will continue to supply power in parallel to the DC buses. and AC fixed frequency services. The APU GCU inhibits APU generator output if external DC power is applied to the aircraft.

The APU electrical load and voltage can be monitored by:
• Selecting GEN 2. on the DC SYSTEM indicator panel
• pushing the DC LOAD METER switchlight. on the APU CONTROL panel

If a statter-generator fault is detected the:
• Starter-generator is disconnected from the right main feeder bus
• Main bus tie and the main to secondary bus ties are automatically opened
• ON segment of the GEN switch light goes off
• WRN segment of the GEN switch light goes amber
• APU caution light comes on
• Master CAUTION switch light flashes.

If the starter-generator overheats, the:
• GEN OHT advisory light comes on
• WRN segment of the GEN switch light goes amber
• APU caution light comes on
• master CAUTION switch light flashes
• APU automatically shuts down.

The APU supplies hot bleed air directly to the supply duct upstream of the air conditioning pack. It also supplies continuous bleed air to the pneumatic deicing system when the APU is operating.??? The deice pressure gage will read approximately 6 psi.

Supply to the air conditioning pack is controlled by the BL AIR switch light, on the APU panel. When the switch light is pushed, the OPEN segment will indicate that the APU bleed air supply valve is open. Check valves prevent supply from APU when engine bleed air available


The Electronic Sequence Unit (ESU), in the APU compartment, automatically controls APU start sequencing, malfunction indication and protection. If the ESU detects a fault (high EGT, high oil temp, low oil press), it shuts down the APU and causes the related APU advisory and caution lights to come on.

If the ESU detects an over speed, the:
• FLR amber segment of the APU PWR switch light comes on
• APU caution light comes on
• master CAUTION switch light flashes
• APU is automatically shutdown.

The over speed protection system can be tested manually by pushing the OVERSPEED TEST switch light, on the APU panel.

An APU start requires either aircraft batteries or external power. During APU spool up, the ESU energizes the APU ignition system and opens the fuel solenoid valve at the proper engine RPM. The ESU then ensures correct acceleration and stabilization of the APU, before arming the bleed air and generator systems for operation.

During APU start, if the APU RUN and STARTER lights come on at the same time, showing an APU starter failure, the right main bus must be de-energized to reduce subsequent damage to the starter. See Vol. 1, Section 2.

If the APU fails to start or accelerate, restart is prevented for two minutes by a timed relay. The APU PWR switch light must be reselected after an APU:
• Automatic shutdown
• Failure to start.

Before shutting down the APU, close the bleed air valve and select the generator off. To shut down the APU, push the APU PWR switch light off.

There is an automatic APU blower system for oil and rear fuselage bay cooling that operates when APU power is on. The blower fan draws air from a spring-loaded door in the maintenance access hatch. The air goes through the oil cooler and is then removed through the ventilation exhaust duct.

The rear compartment has two temperature sensors. If an overheat condition is sensed, the:
• APU automatically shuts down
• RBY OHT segment (amber) comes on
• APU caution light comes on
• master CAUTION switch light flashes.

A fully automatic fire detection/extinguishing system (independent from the engine system) is supplied for the APU. It provides extinguishing capability for APU internal or compartment fire. The system monitors the APU hot section and exhaust when the right essential 28 VDC bus is energized. The system includes:
• A stainless steel fire extinguisher bottle
• A loop sensor routed along the exhaust and shroud
• Control circuit.

If a fire or overheat condition is sensed by the detection loop, the:
• FIRE warning light (red) on the FPP comes on
• CHECK FIRE DET warning light flashes
• master WARNING switch light flashes
• APU caution light comes on
• master CAUTION switch light flashes
• BTL advisory light (amber); then (out) if the bottle has discharged
• APU FUEL VALVE CLOSED advisory light (white) comes on
• RUN segment of APU PWR switch light off.

If a fire is detected, the ESU automatically shuts down the APU and discharges the fire extinguisher. If automatic fire extinguisher discharge fails, the BTL light stays on, the guarded EXTG switch on the APU FPP allows manual discharge.

Once the bottle has been discharged, restarting the APU is prevented until the bottle has been replaced.


Two fire extinguisher discharge indicator discs are located externally at the right rear fuselage. They provide a visual indication of extinguishant discharge. The yellow disc indicates bottle discharge and the red disc thermal discharge if extinguishant escapes to the atmosphere due to thermal expansion.

The Dash 8 Series Q300 Aircraft has an array of standard and optional flight instruments for day and night VFR and IFR operations.

A full set of primary flight instruments are arranged on the pilot's and copilot's instrument panels.
Each panel has:
• Airspeed indicator
• Altimeter
• Inertial Vertical Speed Indicator (IVSI)
• Electronic Attitude Director Indicator (EADI) with inclinometer
• Electronic Horizontal Situation Indicator (EHSI)
• Radio Magnetic Indicator (RMI).
An electromechanical ADI and HSI may be installed in place of the EADI and EHSI

Additional instruments located on the instrument panels and glareshield include:
• Standby attitude indicator
• Standby altimeter
• Radio Altimeter
• Two digital clocks.

The following systems supply critical data to the instruments:
• Air Data System (ADS)
• Attitude Heading Reference System (AHRS).

The term Attitude Director Indicator (ADI) will be used to imply both ADI and Electronic ADI (EADI) and the tern} Horizontal Situation Indicator (HSI) will be used to imply both HSI and Electronic HSI (EHSI) unless stated otherwise.

The Air Data System (ADS) supplies pitot pressure to the:
• Air Data Computer No.1 (ADC 1) and No.2 (ADC 2)
• Pilot's Airspeed Indicator (ASI)
• Copilot's Airspeed Indicator (ASI)

ADC 1 provides aircraft altitude to the pilot's main altimeter and the ADC 2 provides aircraft altitude to the copilot's main altimeter.

There is one pitot head on each side of the aircraft nose. The left pitot head supplies dynamic air pressure to the ADC 1 and the pilot's ASI. The right pitot head supplies the ADC 2 and the copilot's ASI.

The ADS supplies static pressure to:
• ADC 1 and ADC 2
• Pilot's and copilot's ASI
• Pilot's Inertial Vertical Speed Indicator (IVSI)
• Copilot's IVSI
• Standby Altimeter
• Cabin Altitude Controller

There is one static vent plate on each side of the fuselage above and forward of the pitot head. Each static vent plate has three static ports. Each port is cross-vented to the port on the opposite side of the airplane.

One port supplies ADC 1, pilot's ASI and IVSI and the Standby Altimeter. One port supplies ADC 2, copilot's ASI and IVSI. The third port supplies the cabin altitude controller and acts as an alternate source for either of the other two 'should they become blocked. When either STATIC SOURCE selector is set from NORMAL to ALTERNATE, the flight instruments for that side are supplied with alternate static air pressure.

The pitot heads and the static vent plates are electrically heated to prevent icing. Anti-icing heat is controlled by the PITOT STATIC heat switches on the ice protection panel.

The airspeed indicators sense pitot and static air pressure to show indicated airspeed (IAS), and static pressure to indicate maximum operating speed (VMO).

Two inertial vertical speed indicators (IVSI), one located on each instrument panel, sense the rate of change of the static pressure. Aircraft with Traffic Alert and Collision Avoidance System (TCAS) will have the TCAS type IVSI installed.

Two ADCs receive pitot-static pressures and air temperature inputs to independently calculate standard air data functions. Static air temperatures are obtained from two temperature probe inputs. Two advisory display units show static air temperature (SAT) continuously and True Airspeed (TAS) when selected.
Computations and conversions based on altitude, airspeed. and temperature provide the following ADC outputs:
• Indicated airspeed
• True airspeed
• Pressure altitude
• Barometric correction
• Barometrically corrected altitude
• Vertical speed
• Static source error correction
• Digitized pressure altitude for mode C (transponder)
• Overspeed warning

ADC outputs are available to:
• Flight Guidance Computers (FGCs)
• Attitude and Heading Reference Systems (AHRS)
• Advisory displays
• Primary Altimeters
• Autopilot

Both ADCs are operating when power is applied to their electrical buses. Data from one ADC is used for flight guidance computer operation and advisory display temperature presentation. The ADC to be used for temperature information is selected by the HSI SEL switch, on the flight guidance controller panel.
When either ADC senses that the calculated indicated aircraft speed is above the calculated maximum operating speed (VMO), a flight compartment warning horn sounds an intermittent tone. The tone may not sound until the airspeed pointer is as much as 6 KIAS above the VMO pointer. The overspeed warning circuits can be tested by selecting the ADC test switch on the pilot's side panel.

The pilot and copilot primary servo altimeters give a counter drum and pointer display of barometrically corrected altitude from their related ADC. If an ADC fails. the related side altimeter will stop receiving valid altitude data. This will be indicated by a red failure warning flag that will also appear if power is removed from the instrument, or during an ADC test.

An altitude alert system is used to notify the crew when approaching or departing a preselected altitude as selected on the ALTITUDE PRESELECTER. The Altitude selector indicator displays dashed lines if the active ADC fails. Actual aircraft altitude is compared with the preselected altitude on the controller. Visual and aural alerts (optional) are provided.

When in acquisition mode and 1,000 feet above or below the preselected altitude, an optional tone is momentarily sounded when the servo altimeter alert light comes on. When the aircraft is less than 250 feet of the preselected altitude, the alert light goes out and the system switches to deviation mode. If a deviation of more than 250 feet occurs from the preselected altitude, the alert light turns on and the tone sounds momentarily. The deviation alert is the same as the acquisition alert. When the aircraft returns to less than 250 feet of the selected altitude, the light turns off. lf the aircraft does not return to the selected altitude, the preselected altitude must be reset for the altitude alert light to go out.

The Attitude/Heading Reference System (AHRS) (Figure 4F-41) has:
• Two AHRS controllers
• Two Attitude/Heading Reference Units (AHRU)
• Two remote flux valves

The AHRS supplies aircraft attitude and heading data to the:
• ADI, HSI and RMI
• Radio Magnetic Indicators (RMI)
• Flight guidance computers (FGC)
• Weather radar

The AHRU uses vertical and directional gyros and accelerometers to sense rate of aircraft movement, which is then provided to the pilot as pitch, roll, and heading information. The AHRS operating modes are NORMAL for attitude and SLAVED for heading. These modes automatically initialize, if all system components and signals are correct during power-up. In NORMAL mode, ADC true airspeed data adjusts for acceleration-induced errors. The SLAVE mode receives magnetic heading reference from the flux valves.

The RMIs display slaved heading information and are cross-tied to the opposite HSI. The pilot's RMI displays the same heading information as the copilot HSI and vice versa. In SLAVED mode the difference between the indicated and flux valve heading is shown on the slave error indicator of the HSI.

Two reduced-performance modes are also available:
• BASIC mode for attitude
• DG (directional gyro) for heading

BASIC mode, shown in green on the AHRS controller, is entered, automatically when ADC true airspeed data becomes invalid. Operation in BASIC mode will cause decreased attitude accuracy, and the ADI will behave somewhat like a conventional mechanical gimbaled gyro.

In the SLAVED mode, a loss of valid flux valve data will result in a HDG flag on the HSI and a red SLAVE light on the AHRS controller. The pilot's response is to select DG mode. The HDG flag will clear upon entry to the DG mode, but the SLAVE annunciation will remain. The DG mode turns off automatic slaving of the heading, therefore the HSI will operate similar to a free directional gyro.

Approximately 3 minutes is necessary to initialize the AHRS after power is applied. During initialization the aircraft must remain stationary, however passenger and baggage loading, engine start and run-up can be performed. If the AHRS does not initialize after 3 minutes the system must be reinitialized. Pulling and resetting all applicable circuit breakers starts initialization.


To reset the AHRS, it is necessary that both breakers (primary and auxiliary) be pulled out. Resetting each breaker individually will not reset the AHRS.

The initialization time remaining may be monitored by pushing the VG ERECT button on the AHRS controller.

Initialization during flight is not recommended. If it is necessary to do an initialization in flight, the aircraft must be kept in wings-level unaccelerated flight during the alignment.

A 5-second self-test may be started manually at any time, on the ground or in flight by momentarily pushing the TEST pushbutton on the AHRS controller.


This description of the APIRS shows how it operates with the Honeywell SPZ-8000 Integrated Flight Control System.

The dual APIRS installation is the primary attitude and heading reference source. Several operational modes maintain attitude and heading in the event of certain system failures. The standard system consists of the following: dual attitude and heading reference units (AHRU), and the dual flux valves.
• Dual attitude and heading reference units (AHRU)
• Dual attitude and heading control panel (AHCP)
• Dual flux valves.

Pitch, roll, and heading are output to the standard electromechanical or optional EFIS, and the automatic flight control system (AFCS). Attitude and heading data is supplied to other aircraft systems such as weather radar antennas and radio magnetic indicators.

The standard APIRS operating modes are the NORMAL mode for attitude and the SLAVED mode for heading. Following initialization, the APIRS enters these modes automatically, if all system components and signals are valid. In the NORMAL mode, true airspeed from the air data computer is used to compensate for acceleration-induced errors nonnally encountered in a vertical gyro system. In the SLAVED heading mode, the flux valve is used to establish the magnetic heading reference. System operation in this mode is similar to that of a conventional gyroscopically stabilized magnetic compass.

In the SLAVED mode, a loss of valid flux valve data displays a heading (HDG) flag on the heading instruments and the SLAVE annunciator on the AHCP controller lights. The HDG flag clears when the directional gyro (DG) mode is subsequently entered, but the SLAVE annunciator remains lit.

When the DG mode is exited, the APIRS performs an automatic synchronization of the heading outputs to the present flux valve magnetic heading. This feature can also be used if a heading error should develop, while in the SLAVED mode. The error can be removed by momentarily entering the DG mode and returning to the SLAVED mode. This is done by pushing the HDG/DG button on the APIRS controller twice.

In the SLAVED mode, the difference between the indicated heading and the flux valve heading is displayed on the slave error indicator (heading sync indicator) located on the HSI. The card has two symbols: a cross (+) and a dot ('). During straight and level flight, the indicator is generally centered with excursions toward the cross or dot occurring over a 20 to 30 second time period. This activity is normal and indicates good magnetic heading data. In turns, the display may show a steady dot or cross. Following return to straight and level flight, the indicator returns to the centered condition within 2 minutes.

The APIRS has two power source inputs. On the pilot's side, the primary power input is from the left essential power bus. and the auxiliary power input is from the right essential power bus. On the copilot's side. primary power comes from the right main bus, and the auxiliary power from the left: essential bus. Separate circuit breakers control each of these power circuits APIRS shutdown in flight due to power load or bus switching transients is prevented by automatically transferring power within the APIRS to the auxiliary input. When primary power input is restored, the APIRS switches back to the primary power source.

TWO reduced performance modes are also available. The BASIC mode is used for attitude. and DG mode is used for heading. The BASIC mode is annunciated and is entered automatically when true airspeed from the ADC is invalid. The BASIC mode is annunciated and is entered automatically when the pilot's side TAS from the ADC is invalid, only when the aircraft is on the ground. In the air. if the pilot's side ADC fails, the APIRS switches to the copilot's side ADC and does not annunciate the basic mode. When a dual ADC failure occurs in the air. both AHCPs annunciate the basic mode.

APIRS operation in the BASIC mode results in an attitude system that is similar in behavior to a conventional vertical gyro with pitch and roll erection cutoffs, and is subject to drift and acceleration errors. For this reason, APAS operation in the BASIC mode, results in reduced accuracy in attitude.

The DG mode disables the automatic slaving of the heading outputs. The DG mode can only be entered by momentarily pushing the DG button on the AHCP. When the DG button is released, the DG mode is confirmed by lighting the DG button on the AHCP. APIRS operation in the DG mode results in a heading system that is similar to a free directional gyro, and is subject to drift and turn error. For this reason, APIRS operation in the DG mode results in reduced heading accuracy.

While in the DG mode, the heading card can be manually set to any heading using the DG SLEW buttons on the AHCP. The control is inactive in the SLAVED mode. Two slew speeds are used for each direction. The SLOW slew position operates by momentarily pushing the DG SLEW buttons, and is used for fine heading card adjustments. The FAST slew position operates by push ing the DG SLEW buttons for at least 5 seconds, and is used for large heading card adjustments.

The AHCP, is used to control the APIRS system. The descriptions below identify the switches and knobs on the controller.
• DG Slew Buttons -Push the right button (+) to increase the heading indicated on the heading card. Push the left button (.) to decrease the heading values. In the DG mode, without a SLAVE fail indication on the AHCP, the system can be manually slaved with the DG SLEW buttons and the slave error indicator. This is done by pushing the DG SLEW buttons in the right direction to center the en-or indicator between the (+) and ().

Two slew speeds are used for each direction. The SLOW slew speed is operated by pushing the left or right DG SLEW buttons and is used for fine heading card adjustments. The fast slew speed is operated by pushing and holding the DG SLEW buttons for 5 seconds, and is used to make large heading card corrections. When the button is released, the heading directional update stops.

• DG Button -The basic operation of the DG button is to toggle between the HDG and DG modes of operation. When the DG mode is exited, the APIRS performs an automatic synchronization of the heading outputs to the present flux valve magnetic heading. This feature can also be used if a heading en-or should develop, while in the SLAVED mode. The en-or can be removed by momentarily entering the DG mode and returning to the SLAVED mode. This is done by pushing the DG button on the APIRS controller twice.

The DG mode disables the automatic slaving of the heading outputs. The DG mode can only be entered by momentarily pushing the DG button on the AHCP. When the DG button is released, the DG mode is confirmed by lighting the DG button AHCP. APIRS operation in the DG mode results in a heading system that is similar to a free directional gyro, and is subject to drift and turn error. For this reason, APIRS operation in the DG mode results in reduced heading accuracy.
• SLAVE Annunciator -The SLAVE annunciator lights when the system is operating in the SLAVE mode described above.
• BASIC Annunciator -The BASIC mmunciator lights when the system is operating in the BASIC mode described above.
• ATT/HDG ALIGN Button -The align sequence can be manually initiated anytime (including in flight or during initialization) by momentarily pushing the ATTIHDG ALIGN button on the AHCP.

The APIRS is equipped with automatic self-test, that is performed when power is first applied. The test lasts 5 seconds and it displays the following outputs on the ADI and HSI:
• 10° pitch up
• 20° right wing down
• 060° heading, turning at 1 0/ sec toward North
• North heading, turning at 3°/sec toward east
• All APIRS controller annunciators ON
• ATT flag valid for 2.5 seconds, then invalid
• HDG flag valid for 2.5 seconds, then invalid
• Rate-of-turn indicator standard rate (3°/sec) right tum (on optional EFIS EADI).
The flags remain invalid until initialization is complete.

The APIRS system requires approximately 60 seconds to initialize following application of power. The initialization is complete when the ATT and HDG flags clear on the ADI and HSI. During the initialization, the aircraft must remain stationary. Wind gusts and aircraft buffeting are not limiting in this respect. All normal preflight operations, including engine starts and passenger loading, can be carried out while the APIRS is initializing. If the initialization requires more than 60 seconds, the APIRS may have detected excessive aircraft motion. If aircraft movement has occurred during initialization, the APIRS must be recy-cled and a new initialization started. The HSI heading card slews to approximately 60°. The heading decreases at the rate of 1 ° /sec until the heading card indicates north (000°). At this time, the 60 second initialization period is complete and all indications return to normal.

If the heading card stops and does not step to an indication of 000°, the initialization of that APIRS has not been completed satisfactorily. The main and auxiliary DC power to that APIRS should be removed by opening the circuit breakers and then reapplying them to restart the initialization.

Both breakers (primary and auxiliary) must be pulled out. Resetting each breaker individually does not reset the APIRS.

In order to increase satisfactory ground initialization, the following must be considered:

• The aircraft must remain stationary on the ground until the attitude and heading flags are pulled out of view. Normal passenger and cargo loading, engine start, and engine run-up procedures can be performed during the initialization. Wind buffeting is not limiting in this respect. Taxiing or towing the aircraft during APIRS initialization is prohibited.

• Verify that 60 seconds after power is connected to the dc buses, the attitude and heading flags are out of view. If the timer has stopped, the APIRS is not useable and it should be re-initialized. With the aircraft stationary, push theATTIHDG ALIGN button on theAHCP and release it. If the flags do not pull in after 5 seconds, re-initialize the APIRS.

• Verify APIRS and display functions by watching the APIRS test sequence.

Proper display movement, flag operation, and controller lamp operation must be verified. These tests are performed automatically when power is applied. The pilot can initiate the align sequence at any time including during APIRS initialization, by pushing the ATTIHDG ALIGN button on each AHCP.

• Normal preflight taxi checks of pitch, roll, heading, and rate-of-turn, must be made on each system.

As with any magnetic flux value based heading system, taxiing near areas of large magnetic disturbances can cause the APIRS heading display to show an incorrect heading. In such a case, the aircraft must be in the DG mode while taxiing.


• If the aircraft is moved during API RS initialization, both APIRS must be re-initialized. This is done by pulling all four APIRS circuit breakers and then resetting the four circuit breakers to their normal position.


Take-off with one APIRS in BASIC mode is prohibited,
Takeoff with one APIRS in the DG mode is prohibited,

If a malfunction causes the APIRS to revert to the BASIC mode, it is annunciated on the APIRS controller as BASIC. Normal flight operations can be continued in the BASIC mode subject to the limitations of the Aircraft Flight Manual.

After the BASIC mode has been entered, the pilot must avoid sustained, shallow banked turns of less than 6° (e.g., a constant turn to hold DME arc). In addition, particular attention must be paid to ensure correct aircraft trim.

If a heading flag is observed during a flight and the SLAVE annunciator on the AHCP is lit, the DG mode can be selected by momentarily pushing the DG button.

The APIRS heading must be checked every 5 minutes with reference to a knovm accurate heading source. Errors can be removed by using the DG SLEW buttons on the AHCP to set the heading card to agree with the known reference.

If an abnormal indication appears on the ADI or HSI, correct operation of the APIRS can be confirmed by pushing the ATTIHDG ALIGN button on the AHCP.

The autopilot can disengage when the ATT/HDG ALIGN button is pushed.

During taxi, accelerated flight, or turning maneuvers, small, temporary heading differences can be induced in the compass system due to the pendulous nature of the flux valve. Depending on the magnitude of the heading error of a single APIRS, the HSI HDG MISMATCH message can be posted on the advisory display.

The two methods used to correct a heading mismatch are:
• Establish a steady-state, wings-level flight condition for 2 minutes to wash out the error.
• Establish a steady-state, wings level flight condition and push the AHCP DG button twice.

The most common practice is to push the DG button twice. This method instantaneously synchronizes the flux valve heading regardless of the aircraft's attitude. If the two-push method is used, the aircraft must be in wings level, unaccelerated flight to instantly correct the heading information.

1. If the aircraft is not in level, unaccelerated flight, using the two push method can create a new and potentially larger incorrect heading reference error.
2. The compass systems should be synchronized with the two-push method only while on the ground.
3. A pegged compass synchronization annunciator (full . or + indication) on one of the HSI displays is a good indication of which compass system has an error.

In-air initialization is not recommended. If it becomes necessary to perform an in-air initialization, the APIRS requires approximately 90 seconds to initialize following application of power. The initialization is complete when the ATT and HDG flags clear on the ADI and HSI. During the initialization, the aircraft should be maintained in wings-level unaccelerated flight while the APIRS is initializing. If the initialization requires more than 90 seconds, the APIRS may have detected excessive aircraft motion. If aircraft movement has occurred during initialization, the APIRS must be recycled and a new initialization started. The HSI heading card slews to approximately 90° and the FAST annunciator lights. The heading decreases at the rate of 1° /sec. until the heading card indicates north (000°). At this time, the 90 second initialization period is complete and all indications return to normal.

If the heading card stops. and does not step to an indication of 000°. the initialization of that APIRS has not been completed satisfactorily. The main and auxiliary DC power to that APIRS should be removed by opening the circuit breakers and then reapplying them to restat1 initialization.

Both breakers (primary and auxiliary) must be pulled out. Resetting each breaker individually does not reset the APIRS.

To increase the possibility of a satisfactory initialization, the following must be considered:
• Verify that 90 seconds after power is connected to the DC busses, the attitude and heading flags are out of view. I f the timer has stopped, the APIRS is not useable and should be re-initialized. With the aircraft stationary, push the ATT/HDG ALIGN button on the AHCP and release it. If the flags do not pull after 5 seconds, re-initialize that APIRS.


The Electronic Flight Instrument System (EFIS) has two:
• Electronic Attitude Director Indicators (EADIs)
• Electronic Horizontal Situation Indicators (EHSIs)
• Symbol generators
• EFIS controllers

Two electronic flight display units (EADI and EHSI) are installed one above the other on each instrument panel. The EFIS uses color cathode ray tubes (CRTs) to show attitude and navigational data on the electronic display units. There are two EFIS Display controllers to select the format on the EADI and EHSI. Two symbol generators convert data information from the different sources for presentation on the CRTs. Two INSTRUMENT REMOTE CONTROLLERS (Figure 4F-42) are connected to the symbol generators; one controls the heading bug, the other the course select pointer.

Output data presented on the EADJ and EHSI includes:
• Flight director steering commands
• Altitude data
• Navigational and compass data for flight control
• Advisory data
• Warning flags
• Data flags

The EADI shows the attitude sphere together with lateral and vertical flight guidance command bars (or a single cue command pointer). Command bars show computed steering commands to approach and maintain a desired flight path to which the airplane symbol is flown.

The EHSI shows a "plan view" of the airplane position. The indicator shows airplane displacement relative to navigation source radials, localizers, and glideslope beams. At power-up, the EHSJ shows a full compass display (Figure 4F-13, 4F-14, 4F-15), but can present a partial compass mode when selected through the related EFIS controller. The partial compass mode shows a 900 arc of the heading dial.

When area navigation (RNAV) is selected to the EHSI, and MAP is pushed on the EFIS controller, multiple RNAV waypoints are shown in the partial compass mode. With the multiple waypoint MAP mode, the course select pointer and deviation bar are replaced by a smaller course deviation bar and scale at the bottom of the display (Figure 4F-18).

If a display unit (EADI or EHSI) fails, turning the failed unit off will select the composite mode. The composite mode (Figure 4F-16) combines the data from both display units (EADI and EHSI) onto the remaining operational display on the same side. The display unit showing the composite mode continues operating from its symbol generator and EFIS controller.

If a symbol generator fails, pushing the REVN pushbutton above the EADI  will reproduce the opposite side display. The EFIS controller of the operational symbol generator controls both EADI and EHSI displays during reversionary mode. The REVN pushbuttons have no function unless a symbol generator failure occurs.

For further details of each display, controls and junctions, consult Supplement 19 and the Honeywell Pilot's Manual.

Electro-mechanical instruments may be installed in place of the EFIS system. The Attitude Director Indicator (ADI) would replace the EADI and the Horizontal Situation Indicator (HSI) would replace the EHSI. The EFIS controllers and symbol generators are not required when electro-mechanical instruments are installed.

The electromechanical ADI (Figure 4F-23, 4F-24, 4F-25) combines the attitude sphere display together with lateral and vertical flight guidance command bars (or single cue command pointer). These provide the necessary commands to intercept and maintain a flight path.

The electromechanical HSI shows a "plan view" presentation of the airplane position. (Figure 4F-26, 4F-27) The indicator shows airplane displacement relative to navigation radials, localizers, and glide-slope beams.

Two INSTRUMENT REMOTE CONTROLLERS (Figure 4F-32) interface with the HSls.


A standby magnetic compass is located at the top of the windshield centerpost. The standby compass must be used carefully because compass deviation can vary with the electrical equipment powered in the flight compartment. There is a compass correction card on the flight compartment ceiling above the copilot. Windshield heat may cause irregular and incorrect compass indications.

The standby attitude indicator is gyro stabilized and is powered from the left and right 28VDC essential bus. A striped red OFF flag indicates an invalid display due to inadequate gyro speed or input power failure.

A standby barometric altimeter (Figure 4F-35) supplies an alternate display of barometrical corrected altitude. The standby altimeter operates using static pressure from the pilot's static source, and is independent of the main altimeters.

The radio altimeter system (transmitter/receiver and antenna) supplies:
• Altitude above ground-level (AGL) from 0 to 2,500 feet
• Decision height (DH) selection
• Failure annunciation
• Internal self-test

The HSI shows radio altitude. From 0 to 200 feet the display is in 5-foot increments, and above 200 feet, in 10-foot increments. The display is blank if the radio altitude is above 2,500 feet AGL. If the radio altitude input is invalid or inoperative, dashes are shown.

Radio altitude supplies data to the:
• Ground Proximity Warning System (GPWS)
• Flight Guidance Computers
• Air Data Computers
• Flight Data Recorder

There are two Radio Magnetic Indicators (RMI) one is fitted for the pilot and one for the copilot. The RMI shows aircraft heading and relative radio bearing (Figure 4F-9).

Aircraft heading is shown by a compass card. The compass card on the pilot's RMI indicates AHRS 2 heading, and the copilot's RMl indicates AHRS 1 heading.

The radio bearing is shown by two pointers (No.1 and No.2). The pointers can be set for either Automatic Direction Finding (ADF) or VHF OmniDirectional Range (VOR) bearing indication.

The pointer will park in the horizontal position if the:
• VOR/rr,S system is in the rr,s mode
• VOR is not receiving a signal
• ADF is not receiving a signal

Two digital electronic clocks are located in the glareshield.

Each clock supplies:
• Clock Time (TIME) - UTC or Local.
• Flight Time (FT)
• Elapsed Time (ET)

These functions are shown in hours, minutes, and seconds displayed by setting the clock function select switch to the necessary function. All functions continue to operate regardless of which one is selected for display. The flight time function starts and stops automatically by weight-off-wheels and weight-on-wheels signals from the Proximity Switch Electronic Unit (PSEU).

A separate clock battery powers the internal clock functions, and 28 VDC powers the display.

The avionics compartment is cooled by all avionics cooling fan which draws air across the equipment shelves through extraction ducts and discharges it under the cabin floor. The fan requires 28 VDC and a temperature above a predetermined value to operate.

If the fan fails, cooling by natural convection is adequate for limited operation. If the fan fails, a light above the wardrobe, comes on.

The Dash 8 Series Q300 uses Direct Current (DC) and Alternating Current (AC) power systems. The DC system consists of storage, generation, distribution, and system monitoring. The AC system consists of generation, distribution, and system monitoring. DC and AC power supply electrical power to related aircraft systems.

The DC power system is supplied by two NiCad batteries, two engine driven starter/generators, two Transformer Rectifier Units (TRUs) and an optional Auxiliary Power Unit (APU). The TRUs supply 28 VDC and are powered by two engine driven 115 VAC generators. Fixed-frequency AC power, required for avionics and system indicators, is supplied by three solid-state static inverters and two step-down transformers.

There are both DC and AC external power receptacles for Ground Power Unit (GPU) connection.

Electrical power is distributed by an electrical bus system. It reconfigures for individual power source and bus failures by automatically closing and opening bus tie contactors.

All AC and DC aircraft services can be operated from the AC generators or the AC external power alone.

The entire aircraft electrical schematic is shown in Figure 7-17.

Electrical power sources for the DC system (Figure 7-18) include the:
• Main and Auxiliary Batteries
• Two Starter-Generators
• Two Transformer Rectifiers Units (TRU)
• DC External
• APU Starter/Generator (Optional)

The power sources supply power to the following buses:
• Left (L) and Right (R) Essential
• Left (L) and Right (R) Main
• Left (L) and Right (R) Secondary

Generator output is applied through main feeder buses to the main and essential buses. TRU output is applied through secondary buses.

Setting the lever locked MAIN BUS TIE switch to MAIN BUS TIE. manually connects the lett and right main buses. The buses can also be automatically connected if one or two power sources are not available. without affecting the supply of power to the buses.

The main and auxiliary 24-volt NiCad batteries are in the lower left nose compartment. The main battery has a 40-amp-hour capacity and the auxiliary a 15-amp-hour capacity. An optional 40-amp-hour auxiliary battery can also be installed in the auxiliary position. The batteries are used for engine starting if ground power is not available, and supplies backup power for at least 30 minutes in flight of essential services.

Setting the BATTERY MASTER switch to the BATTERY MASTER position connects both batteries to the essential buses. In addition to the BATTERY MASTER switch, each of the batteries has a dedicated switch. Setting the MAIN BATT switch to MAIN BATT, connects the main battery to the right main bus, and through the bus tie, to the left main bus. Similarly, setting the AUX BATT switch to AUX BATT connects the auxiliary battery to the left main bus for charging. Diodes prevent the auxiliary battery from powering the main buses.

Battery power cannot be applied to the secondary buses.

MAIN BATTERY and AUX BATTERY caution lights come on when the related battery is not connected to its feeder bus.

A BATTERY TEMPERATURE monitor, on the overhead panel (Figure 7-4), uses display sensors to show battery temperatures. If a display sensor fails, one amber light segment, at the 60°C point of the related battery temperature scale, comes on. If an overheat sensor detects a battery temperature of more than 65 °C, the related MAIN BAT HOT or AUX BAT HOT warning light flashes. If an overheat sensor fails, a related SENSOR FAIL light, on the BATTERY TEMPERATURE monitor, comes on.

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 63% NH. It then changes to generator operation. If the DC GEN switch is in the 1 or 2 position, after a successful engine start, the generator will be available. Each generator output is monitored and controlled by its Generator Control Unit (GCU). After engine start, the GCU makes sure the generator supplies 28 VDC (300 Amps max) to its feeder bus, regardless of load.

If the DC generator is disconnected from its feeder bus, the related #1 DC GEN or #2 DC GEN caution light comes on. If a generator overheats, the related #1 DC GEN HOT or #2 DC GEN HOT caution light comes on. The caution light goes off when the temperature drops below limits.

Two Transformer Rectifier Units (TRU) change three-phase, 1l5-volt, variable-frequency AC input power to 28 VDC (200 Amps max) output.

The TRU output voltage can range between 25.0 and 29.5 volts depending on secondary bus load.

The TRUs supply DC power when AC generator or AC external power is available on the related variable frequency AC bus. Under normal conditions, each TRU powers its related secondary bus.

If either TRU is off line or failed, the related L TRU or R TRU caution light comes on. If the related temperature sensor detects an overheat condition, the L TRU HOT or R TRU HOT caution light comes on. The caution light goes off when the temperature drops below limits.

The two TRUs alone can power the entire DC system.

Four bus tie contactors, controlled by DC logic relays, connect the appropriate feeder buses together, when there are one or more DC power sources not operating. !f up to two DC power sources are not operating, the relays makes sure all DC buses are powered.

The four relays connect the following feeder buses:
• L Main to R Main
• L Secondary to R Secondary
• L Main to L Secondary
• R Main to R Secondary

On-line signals from the GCUs and TRUs are used to determine which bus tie(s) to close in order to maintain power on the affected feeder bus, after a power source failure.

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 left and right Main to Secondary bus ties close, and the TRUs supply power to all the DC buses. If both TRUs fail, the left and right Main to Secondary bus ties close, and the DC generators supply power to all the DC buses.

The Bus Bar Protection Unit (BBPU) provides the starter-generators and batteries with automatic overload protection from a bus fault (short circuit).

If a bus fault occurs, the BBPU prevents the two main/secondary bus ties as well as the main bus tie from closing, isolating the bus. As a result, the DC BUS caution light comes on.

If the fault is still present after 7 to 10 seconds, the BBPU will shut off the affected generator, and isolate the related battery from the faulted main feeder bus. The MAIN BATTERY or AUX BATTERY caution light and related #1 DC GEN or #2 DC GEN caution light comes on. The BBPU continues to monitor the operating buses. All main DC services on the faulted bus side will not operate.

Manual operation of the main bus tie through the MAIN BUS TIE switch is not possible once the BBPU has reacted to a fault.

If the fault subsequently clears, or fault generator source is isolated, power may be restored with the BUS FAULT RESET switch (Figure 7-2).

The engines can also be started from battery power. During a battery start, only the main battery is used in the starting process. The auxiliary battery is isolated from the main bus by a diode. After one engine starts, with its generator selected, the generator powers both the left and right main buses. After the second engine starts with its generator selected, both bus ties connecting the main to the secondary buses automatically close, energizing all DC buses.

After an aircraft NiCad battery start, the charge rate for fully serviceable batteries will decrease after approximately three minutes.

A Ground Power Unit (GPU) can be connected to the DC external power receptacle on the left side of the nose. The GPU can supply DC power for the aircra1t DC system, and for engine starting.

With DC external power supplied. and the DC EXT PWR switch set to EXT PWR, the main and secondary DC buses receive power. With DC external power, all DC generator connections to the main buses are inhibited. DC external power also isolates the batteries from the main feeder buses, and the MAIN BATTERY and AUX BATTERY caution lights come on.

Batteries on line with ground power is an available MOD.

Battery temperatures and charging rates must be continuously monitored while charging.

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. Both bus ties. connecting the main to the secondary buses. will stay closed until DC power is available from the TRUs.

If the external voltage is more than 31.5 VDC, external ground power is automatically disconnected from the system. If the external power over voltage is corrected, the external power source can be reselected by setting the DC EXT PWR switch to OFF, and then back to EXT PWR.

The DC power system is monitored from a DC SYSTEM monitor panel. The panel uses four digital displays to monitor voltage and load of the batteries. generators. and TRUs.

A BUS VOLTS display shows the bus voltage of the related essential, main or secondary bus selected by the bus voltage display selector. A MAIN BATT or AUX BATT LOAD display shows the charge/discharge rate of the related battery. A LOAD display shows the electrical load of the related DC generator or TRLJ selected by the power source load display selector.

Fixed-frequency AC power is supplied to the related buses by three solid-state static inverters:

• Primary
• Secondary
• Auxiliary

The buses then supply electrical power to:
• Navigation instruments
• Communication equipment
• Flight instruments
• Certain system indicators

The left essential bus supplies DC power to the primary inverter. The right essential bus supplies DC power to the secondary inverter. The left main bus supplies DC power to the auxiliary inverter. The primary and secondary inverters supply 115 VAC to the left and right 115 VAC buses respectively. The primary, secondary and auxiliary inverter outputs are in parallel. The auxiliary inverter may be selected to supply either 115 VAC bus, through the AUXILIARY INVERTER switch. The 115 VAC buses power the related 26-VAC buses from step-down transformers.

A paralleling control box. operating with the inverter switches. controls the inverter output to the 115 VAC buses. A bus tie circuit breaker connects both 115 VAC buses, so that demands on all buses are shared by the operating inverters.

If an inverter fails. the paralleling control box disconnects the inverter. and the related PRI INV, AUX INV, or SEC INV caution light comes on. The two operating inverters then supply AC power to the L and R 115- VAC buses. If a second inverter fails, the one operating inverter can supply sufficient AC power however, the load must be monitored.

A 26V-bus fault or transformer failure causes the L 26 AC or R 26 AC caution light to come on. If a bus fault occurs on the 115- VAC bus. the bus tie circuit breaker connecting the left: and right IIS-VAC buses opens to isolate the malfunctioning bus. The inverter( s) connected to the malfunctioning bus may subsequently fail if the fault persists. The related 26-VAC transformer and its bus will also lose power.

Alternating- current (AC) generators and inverters supply both variable frequency and fixed-frequency AC power to the electrical system.

A 115 VAC generator, on the related Propeller Reduction Gearbox (RGB), supplies variable frequency AC power to the AC buses. Each generator is controlled by a related Generator Control Unit (GCU) and AC GEN switch, on the overhead panel. Internal logic GCU circuits operate the bus contactors which connect the AC generators to the variable frequency buses. Since the GCUs are DC power from the essential buses, the essential buses must be energized for the AC generators to operate.

AC power is available once the condition levers are in the MIN to MAX range, for :
• Anti-ice heating
• Electric pumps
• Powering the TRU s

If an AC generator fails or is not supplying power, the related #1 AC GEN or #2 AC GEN caution light comes on, and the operating generator supplies the aircraft's AC electrical load. An automatic cross tie function, controlled by the AC GCU logic circuits, makes sure that all variable-frequency buses are powered when only one AC generator is on line. If a fault condition exists, the GCU of the inoperative generator closes the bus contactor. The operating generator will power both AC buses.

The AC generators are protected from bus faults by the GCUs that detect any excessive load that might result from a short circuit on a bus. Once a heavy load is detected, the GCU isolates the bus, and causes the related LAC BUS or R AC BUS caution light to come on.

The # 1 AC GEN HOT or # 2 AC GEN HOT caution light comes on if an AC generator overheats.

A Ground Power Unit (GPU) can be connected to the AC external power receptacle in the right nacelle, to supply AC power to the aircraft. This is used primarily for maintenance purposes.

An external power switch, on the overhead AC CONTROL panel, connects power directly to the left and right variable-frequency buses, supplying power to all AC and DC buses.

A malfunction in the AC external power source will cause it to be disconnected from the aircraft electrical system.

AC power is monitored from an AC SYSTEM monitor panel. The panel uses four digital displays to monitor voltage and load of the three inverters and six variable frequency buses.

An INVERTERS VOLTS display shows the output voltage of the related inverter selected on the inverters display selector. An INVERTERS LOAD display shows the load of the related inverter selected. A VARIABLE FREQUENCY VOLTS display shows the output voltage of the related variable frequency bus selected on the variable frequency display selector. A VARIABLE FREQUENCY LOAD display shows the load of the related variable frequency bus selected.

The fire protection system supplies detection, indication and extinguishing of fire or smoke conditions.

There are fire or smoke detection and extinguishing systems for the engines, optional Auxiliary Power Unit (APU), baggage compartments, and the lavatory. There are portable fire extinguishers in the flight and passenger compartments. Refer to Chapter 3 for a description of APU fire detection and extinguishing.

Engine nacelle fire or overheat is detected by continuous-loop sensors routed around vital areas of the engine. Two sensors are installed to provide additional protection. Two extinguisher bottles provide nacelle fire extinguishing, one for each nacelle, with crossover capability to the opposite nacelle.

Warning of smoke in the baggage compartment is provided by smoke detectors on the ceiling and aft face of the compartment divider bulkhead.

A smoke detector is installed in the lavatory and an extinguisher bottle is located above the disposal bin.

Portable extinguishers are available for use to control smoke or fires in the flight compartment, cabin areas, and baggage compartment.

The aircraft fire detection system senses fire or overheat conditions for engine fire zones. When anyone of these conditions is sensed, the system supplies a visual and aural warning to the flight compartment.

The engine fire detection systems provides visible warnings of fire or overheat conditions in the nacelles. and sensor loop faults. Identical and independent sub-systems for each nacelle are provided. Each consist of:
• a sensor loop routed in parallel throughout the nacelle (Loop A)
• a sensor loop routed in parallel near the intercompressor (Loop B)
• responder units
• fire detector control units
• associated flight compartment warnings and control switches

The sensor loops use tubes filled with helium gas to check for fires or overheat conditions. The alarm and fault signals are provided by two switches, the alarm switch and the integrity switch. The signals are processed by the control unit then sent to the fire protection panel.

The alarm switch monitors the pressure of the loop. A pressure increase causes the Fire signal, a pressure loss causes the related FAULT A or FAULT B advisory light

The detection control unit also warns of failures in the A and B loops wiring.

The FAULT A or FAULT B advisory light comes on if:
• pressure loss detected in any loop
• power loss to the alarm and integrity switches
• A defective control unit or responder

A rotary loop selector for each engine selects the operating loop(s). The selector is normally left at BOTH (loop A and B armed). If a fault is detected in either loop system, the related FAULT A or FAULT 8 light comes on. The CHECK FIRE DET warning light flashes. Turning the LOOP SELECTION selector to the opposite (8 or A) loop disarms the defective loop. The CHECK FIRE DET warning light and FAULT light go off. The remaining loop continues to provide fire detection to the nacelle.

When detection has been deactivated, the remaining loop's failure monitoring circuit is prevented from warning of a power supply failure to that loop.

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.

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 discharging the first bottle does not put out the fire, the second bottle can be discharged. Bottle discharge is controlled by EXTG switches on the FPP.

Two extinguisher discharge indicator discs are located below the left inboard flap on the exterior of the aircraft. If the red disc is missing, the extinguisher has vented to the atmosphere, due to thermal expansion. If the yellow disc is missing, a bottle has discharged into an engine nacelle.

If an engine nacelle fire or overheat condition occurs, the indications:
• Both ENGINE FIRE PRESS TO RESET switchlights (red) flash
• CHECK FIRE DET warning light (red) flashes
• MASTER WARNING light (red) flashes
• PULL FUEL OFF tire handle (red) comes on
• Fire alarm bell sounds (optional)

Pushing either ENGINE FIRE PRESS TO RESET switch light cancels the light and silences the bell.

When an engine fire is detected and the PULL FUEL OFF handle is pulled, the extinguisher arm lights turn on to indicate that the control circuits are armed. and the bottles are fully charged. Pulling the handle also closes the emergency fuel and hydraulic shutoff valves to the Engine Driven Pump (EDP) of the respective engine. Electrical power for the fire extinguishing system is direct from the hot battery power bus through the extinguisher discharge switch.

When discharge of the extinguisher is initiated. it flows until the bottle is empty, after which, the associated EXTG amber light goes out.

Smoke detection provides visual flight compartment warning of smoke or fire in the baggage area. The baggage compartment has two smoke detectors, one on the ceiling and one on the bulkhead. If either smoke detector senses smoke, both the SMOKE warning light and the MASTER WARNING switchlight flash. Resetting the MASTER WARNING switchlight will cause the SMOKE light to come on steady for as long as the smoke is present. A BAGGAGE SMOKE WARNING test switch, on the FPP, is used to test each detector circuit.

There are three portable fire extinguishers, one in the flight compartment, and two 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.


If a fire extinguisher is to be discharged in the flight compartment, all crew members must wear their oxygen masks, with the EMERGENCY position selected (100% oxygen at a positive pressure).

There is one smoke detector in the lavatory compartment mounted on the top forward lavatory wall above the waste bin. If the smoke detector senses smoke. the:
• Detector audible warning tone sounds
• Detector alarm light (red) comes on
• LAVATORY SMOKE (red) annunciator light. on the two Advisory Light Panels. comes on
• Chime (high) sounds in the cabin speakers (PA system).

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. Confirm the green power indicator light is on. 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 LAVATORY SMOKE annunciator lights go off

The lavatory waste bin compal1ment is protected by a thermally activated fire bottle with no electrical interface. If a fire occurs in the waste bin, and the temperature of the end caps of the tubes increases to a set point, they melt. The suppressant is then discharged into the bin.

The Dash 8 Series Q300 primary flight controls consist of ailerons and roll spoilers, rudder, and elevators. Secondary flight controls consist of flaps.

All flight controls may be operated from either the pilot or copilot position. The primary flight controls consist of ailerons and spoilers for roll control, elevators for pitch control and a two-sectioned (fore and trailing) rudder for yaw control. Roll spoilers and fore rudder are hydraulically powered.

Secondary flight controls consist of wing flaps and a flight control gust lock mechanism.

Trim position indicators, the POWERED FLIGHT CONTROL SURFACE (PFCS) indicator and a series of caution lights, provide indications for flight control positions or malfunctions.

Yaw control is provided by a two section rudder. The rudder is powered by the No. I and No.2 hydraulic systems and controlled through the movement of either the pilot's or copilot's rudder pedals. A yaw damper operates through the rudder control system to improve directional stability. A summing unit receives inputs from the trim, rudder pedals, and yaw damper systems to move the rudder as required. An artificial feel unit provides the pilot with feedback pressure for the rudder.

The rudder has two sections, fore and trailing. The fore rudder is hinged to the vertical stabilizer and is operated by an upper and a lower actuator. The No. 1 hydraulic system powers the lower actuator and the NO.2 hydraulic system powers the upper actuator.

The trailing rudder is hinged to the fore rudder and deflects twice as far as the fore rudder. Rudder deflection is shown on the PFCS indicator. With 0 flap selected, maximum rudder deflection is 12° left or right to the limit rudder travel. With flaps selected to 5°. 10°. 15°. or 35°, fore rudder deflection is 16° left or 18° right.

A pressure regulator controls hydraulic pressure to the rudder actuators. At airspeeds greater than 140 KIAS. the air data computers (ADCs) signal the regulator to reduce hydraulic pressure to 900 psi. At airspeeds less than 140IAS. hydraulic pressure is increased to 1,500 psi. If the No. I or NO.2 hydraulic system fails. hydraulic pressure to the operating rudder actuator increases to 3000 psi regardless of airspeed. This turns on the RUD FULL PRESS caution light. If hyclraulic pressure fails to change over as scheduled. the RUD PRESS caution light comes on.

If an actuator jam occurs, the RUD I or RUD 2 PUSH OFF switchlight on the glareshield panel will come on. The RUD I or RUD 2 PUSHOFF switchlight must be pushed and will depressurize the affected actuator. The # 1 RUD HYD or #2 RUD HYD caution light will come on when the pressure drops in the affected system.

Rudder trim is maintained by biasing the neutral position of the fore rudder. This is accomplished by an electric actuator which is controlled by the RUDDER TRIM: switch. The actuator provides two speeds of operation. Selecting the RUDDER TRIM switch left or right of center, trims the rudder at a slow rate allowing for small trim changes. A selection to the full L or R position provides a faster rate of trim for larger trim changes. Amount of trim is displayed on the RUDDER TRIM: position indicator, and rudder position is displayed on the PFCS indicator.

A yaw damper system operates independently or in conjunction with the Automatic Flight Control System (AFCS). It improves directional stability and turn coordination by providing compensating inputs to the rudder whenever the aircraft deviates from coordinated flight. Coordinated turns are maintained by repositioning the fore rudder up to 5° left or right of an existing rudder setting.

The yaw damper is engaged by a switchlight located on the Flight Guidance Controller (FGC) and labeled YD. The yaw damper should be selected off prior to takeoff.

Pitch control is maintained by two mechanically controlled elevators. The elevators are attached to the trailing edge of the left and right horizontal stabilizers. The left control column operates the left elevator and the right control column operates the right elevator. However, both control columns are connected to each other by the pitch disconnect system, so that they operate together.

A trim tab and a spring tab are located on each elevator surface. Elevator trim is normally done manually.

There are no caution lights associated with pitch control. There are no indicators for elevator position in the flight compartment.

A spring tab is located on the inboard trailing edge of the left and right elevator. It provides an aerodynamic assistance to elevator movement to reduce elevator control forces.

Pitch trim is controlled by a single trim tab on the left and right elevator trailing edge outboard of the spring tabs. The ELEVATOR TRIM wheels in the flight compartment manually control the trim tabs.

A trim tab position pointer provides indication of trim tab position. The pointer moves between ND (nose down) and NU (nose up) positions with trim wheel operation. A TRIM LT switch on the console controls the light for the trim tab indication window.

If failure of the manual elevator trim occurs. the standby elevator trim can be used to control the aircraft in pitch. The standby elevator trim system is electrically operated and controlled by the STBY ELEVATOR TRIM switches. on the pilot's or copilot's side panel.

The Auto Flight control System (AFCS) normallylly controls the elevator trim servo. When the STBY ELEVATOR TRIM switch is set to ARM. this deactivates the AFCS control from the elevator trim servo. The STBY ELEVATOR TRIM switches now control the elevator trim servo. The manual elevator trim wheels rotate when STBY ELEVATOR TRIM is selected to NOSE DOWN or NOSE UP.

The left and right control columns are mechanically connected to each other through the pitch disconnect mechanism. If a pitch jam occurs in either control circuit, the two control columns can be disconnected from each other by pulling the PITCH disconnect handle on the left side of the centre console. When the handle is in the engaged position. the pilot's and copilot's control columns are connected to each other by a clutch.

When the handle is pulled out and rotated 90deg. the clutch disengages and the two control columns. The pilot with the free control column will have pitch control.

With the PITCH disconnect handle pulled. the autopilot must be disengaged.

Roll control surfaces consist of cable-operated ailerons assisted by hydraulically powered roll spoilers. The pilot's wheel controls the roll spoilers, while the copilot's wheel controls the ailerons. The left and right control wheels are mechanically connected so that rotating either control wheel operates ailerons and spoilers to provide roll control. In the event of a roll control jam, the left control column can be separated from the right control column by pulling the roll disconnect handle. This separates the roll spoilers from the ailerons and allows the pilot with the unjammed wheel to maintain roll control.

An aileron is located on the outboard end of each wing. A geared tab provided on the left aileron and a geared/trim tab on the right aileron assist the pilot in roll control.

An electrically controlled trim tab on the right wing provides aileron trim. Aileron trim is accomplished by pushing the AILERON TRIM rocker switch to reposition the tab. An AILERON TRIM position indicator, adjacent to the trim switch, shows aileron trim position  [Left or Right wing down (LWD or RWD)].

The spoilers operate in proportion to the up going aileron to provide roll control. Turning either control wheel normally operates both the spoilers and ailerons. No. I hydraulic system powers the inboard spoilers and No.2 hydraulic system powers the outboard spoilers. At airspeeds more than 140 IAS, the outboard spoilers are disabled.

During flight, if a jam occurs to the aileron cable linkage, spoiler cable linkage or actuator linkage, will result in a roll control jam. If an actuator jam occurs. applying a slightly increased roll force to the control wheel disengages a clutch linkage in the spoiler system to regain roll control. This causes either SPLR 1 or SPLR 2 PUSH OFF switchlight to come on indicating a jam.

If the PUSH OFF switchlight stays on continuously while attempting to maintain wings level, a spoiler may be jammed in the extended position. The related switch light must be pushed off and will depressurize its system causing the associated ROLL SPLR INBD HYD (all airspeeds) or ROLL SPLR OUTBD HYD (below 140 IAS) caution light to come on. During subsequent roll control inputs, control wheel movement will not be as smooth, due to the clutch engaging and disengaging.

If an aileron or spoiler cable linkage jam occurs, further roll inputs will be prevented. To regain roll control. the ROLL DISC handle must be pulled and turned 90° to determine which pilot has roll control. The pilot with the unjammed wheel flies.

Left Control Wheel Free
If the left control wheel is free, only 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 deflection required to maintain level flight exceeds 50deg. displacement from neutral. SPLR I and SPLR 2 switchlights will come on. This would indicate to one or both spoilers deployed on the same side.

If the SPLR I and SPLR 2 switchlights remain on continuously, both must be pushed off. This will depressurize the affected spoiler system and turn ON the ROLL SPLR INBO HYD (all airspeeds) and ROLL SPLR OUTBO HYD (below 140 KIAS) caution lights. Any spoiler that may have been jammed in the extended position will retract. Roll spoiler positions are presented on the PFCS indicator.

With the ROLL disconnect handle pulled. the autopilot must be disengaged.

Two roll spoiler switchlights, labeled INBO and OUTBO. provide an alternate method of deactivating the roll spoilers. These switchlights perform the same function as the SPLR 1 and SPLR 2 switchlights on the glareshield panel. They are refered to in the ROLL SPLR INBO HYO or ROLL SPLR OUTBO HYO caution light checklist.

Single-slotted inboard and outboard fowler flaps are attached to the trailing edge of each wing. The flaps are connected to screw jacks which are operated by a primary drive shaft. A Flap Power Unit (FPU) actuated by the flap selector, operates the flap drive system and moves the flaps to their selected positions.

The flap lever quadrant provides gates at 0°, 5°, 10°, 15° and 35° positions. These positions are also shown on the FLAP position indicator.

A trigger on the FLAPS lever must be raised out of the gate before the selector can be moved from one flap position to the next. When a flap selection is made, the flap trigger must be released before reaching the intended position.

Microswitches in the flap selector quadrant are actuated by the FLAPS selector lever. Signals from the microswitches are sent to the Ground Proximity Waming System (GPWS), autopilot, stall warning systems, and No.1 and No.2 standby hydraulic pumps. The FLAP position indicator also sends signals to the stall warning computer and take off warning horn. If the flaps are selected less than 5°, or more than 20° during take off, the take off warning horn will sound. The T/O FLAP switchlight, on the pilot's side console, must be pushed to over-ride the warning horn.

The FPU is a hydraulic motor, which receives inputs from the No. 1 hydraulic system and electrical inputs from the flap lever trigger.

The primary drive system operated by the FPU extends and retracts the flaps. When flap selections are made, rotation of the primary drive positions the flaps at the selected position. Primary drive rotation stops when the flaps reach their selected position.

The secondary drive consists of a flexible shaft coupled to the primary drive system through two transfer gearboxes. Normally they operate together, but only the primary drive is under load.

If the primary drive breaks, torque is transferred to the secondary shaft, then back to the pri111my shaft outboard of the failure. This prevents a split flap condition. Flap movement continues to the selected position. When the secondary drive operats under load, the FLAP DRIVE caution light comes on.

When the flap lever is in the 0° position, all flaps are retracted, and the flap position indicator points to 0° . Squeezing the flap lever trigger energizes the FPU. Moving the flap lever to 5°, 10°, 15°, or 35°, signals the FPU to extend flaps to the selected position. The FPU remains energized for 70 seconds to allow sufficient time for flap movement to the selected position.

During flap selections, to avoid excessive loading of the flap drive system, flaps must be allowed to come to a complete stop before the selector lever is reversed.

If a flap jam occurs during extension or retraction, the flap system automatically shuts down. Flap operation will continue if flaps are selected in the reversed direction away from the jam.

The FLAP POWER caution light comes on if hydraulic pressure in the flap power unit is low during flap extension and retraction.

The FLAP DRIVE caution light comes on if the secondary flap drive system comes under load. The flaps will continue to operate and may be operated normally for the remainder of the flight.

A 0° flap landing will be required with either a loss of No. 1 hydraulic system fluid or a total flap failure. To override the Ground Proximity Warning System (GPWS), TOO LOW FLAP call on approach, push the GPWS FLAP OVERRIDE switchlight, on the pilot's side panel

Ailerons and elevators are locked with the control lock lever. The elevators lock in the nosedown position, and the ailerons lock in the neutral position. The gust lock system is operated by a CONTROL LOCK lever on the power quadrant, forward of the POWER levers. 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. If a gust lock cable fails, the gust lock mechanism fails safe, - - to the unlocked position.

To unlock the 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 Control System is a hydraulically powered flight control system. Gust lock protection for the rudder system is provided by retained hydraulic pressure in the rudder actuators. The retained pressure provides stiffness to the rudder surface, which resists movement from wind forces.

Fuel is contained in two integral main wing tanks designated No.1 (Left) and No.2 (Right). The fuel system provides for indicating, storing, venting, fuel feeding and scavenging, refueling/defueling, and transfer. Only tank to tank transfer is available; there is no engine crossfeed capability. The aircraft may be pressure or gravity refueled.

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 within 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 available through two overwing fuel filler ports.

Fuel Quantity indication is measured by 6 capacitance type fuel probes in each tank. The fuel quantity of each tank is shown on related FUEL QTY gauges, on the FUEL CONTROL panel. Pushing and holding a momentary action QTY TEST pushbutton, on the FUEL CONTROL panel, cause both gauges to indicate full.

If the QTY TEST pushbutton is pushed during refueling, the refueling will be interrupted.

Two magnetic dipsticks (magnasticks) 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 magnsticks are calibrated rods with a magnet attached to the top. 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. Due to wing dihedral, two dipsticks are required for each tank. If the inner dipstick does not drop when it is released, the fuel level is at the top of the tank at that location. Total fuel quantity must then be read from the outer dipstick. In either case, fuel quantity must only be read from one magnastick.

The fuel temperature in the left collector tank is shown on a FUEL TANK TEMP gauge. Each engine's inlet fuel temperature, after it has passed through the Fuel Oil Heat Exchanger (FOHE), is shown on the related FUEL TEMP gauge.

There are two integral (wet) wing tanks that extend laterally from the nacelle to the rib just inboard of the ailerons. Two auxiliary tanks are optional. 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 nacelle and collector bay. The collector bay is located at the inboard part of the wing tank. Fuel is contained in the main tanks and the collector bays. Water drain valves are located at the low points of the tanks, surge and collector bays.

Total useable fuel from both tanks is 5,678 Ib (2 574 kg). The maximum lateral imbalance permitted between tanks is 600 pounds (272 kg).

The surge bay is used for main tank venting and fuel recovery. One outboard float vent valve and one inboard vent line, control the pressure between the related surge bay and main tank. The float vent valve, located near the top of the tank, opens and closes 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. This system fills the collector bay faster than it can be pumped out by engine feed or fuel transfer, therefore the excess is returned to the main tank.

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 the related collector bay level drops below approximately 130 Ibs (59 kg).

Fuel to each engine is fed from the collector tank, from a primary ejector pump or an AC driven auxiliary pump, and delivered to the engine driven pump. From there it is delivered to the Mechanical Fuel Control (MFC) before being sent on to the engine. If the engine driven pump inlet pressure drops below limits, the related #1 or #2 ENG FUEL PRESS caution light comes on.

An 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 TANK 1 or TANK 2 AUX PUMP switches, on the FUEL CONTROL panel, control the auxiliary pumps manually. A green AUX PUMP advisory light comes on when the related pump is providing boost pressure.

The engine feed shutoff valve closes when the related PULL FUEL OFF handle is pulled. Green and white FUEL VALVE lights, on the Fire Protection Panel (FPP), show when the valve is open or closed.

The fuel is filtered and heated by the FOHE before entering the engine pump, If the fuel filter is becoming blocked( increased back pressure), a related (#1 or #2) FUEL FLTR BYPASS caution light comes on. Fuel can bypass the filter for continued operation.


Fuel can be transferred from one tank to the other to correct fuel imbalances or for fuel management.

A lever locked TRANSFER switch, on the FUEL CONTROL panel, controls the 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 AUX PUMP advisory light to turn green, Electrically operated fuel transfer shutoff valves open for fuel transfer and close when the transfer is stopped. The float control valve of the receiver tank opens to allow fuel transfer.

Once selected, fuel transfer will continue until deselected. If fuel transfer is not deselected a protective Pilot Valve will close thereby stopping the fuel from being pumped overboard.

Fuel cannot be transferred if the FUELING ON caution light is on. The FUELING ON caution light is on if the refuelldefuel master switch is selected to REFUEL or DEFUEL.

All refuel/defuel operations are controlled by the REFUELIDEFUEL panel. Access to the panel is gained through a flush door on the rear underside of No.2 nacelle. DC power must be available for refueling.

When the Master refuel/defuel switch is selected to REFUEL, the POWER ON light turns on. The FUELING ON caution light also comes on when the refuel/defuel master selector is at the REFUEL or DEFUEL position. The two FUEL QUANTITY CONTROL AND REPEATER indicators are repeaters of the MASTER FUEL QUANTITY indicators in the flight compartment.

The REFUELIDEFUEL panel access door can be closed with the master switch in the REFUEL or DEFUEL position. The only indication in the flight compartment that REFUEL or DEFUEL is still selected is the FUELING ON caution light being on. Fuel cannot be transferred when the FUELING ON caution light is on.

When a REFUEL selection is made, the master refuel/defuel valve opens (MASTER VALVE CLOSED light out), and with fuel pressure the dump valve opens (DUMP VALVE lights on). If the tanks are to be refueled, the desired quantity is selected by positioning the bugs with a set knob on the face of each gauge. When the selected quantity is reached, the control valves automatically shut off fuel flow to the tanks.

If fuel does not automatically shut off, the protective Pilot Valve will cause the flow control valve to close and stop the fueling then the tank is full.

When fueling is in progress, advisory lights on the REFUELIDEFUEL CONTROL panel show the position of the dump valves. A dump valve in each main tank protects the tank from damage due to over pressure. If the fuel tank over fills, excess 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.

The refueling control circuits are armed through dump valve inputs to ensure the dump valves are open before refueling can take place.

A pre check test system simulates full tanks thereby ensuring operation of the overfill shutoff system. With the switch in the PRE CHECK TEST position a full tank is simulated for the related tank, and the pilot valve shuts off refueling to that tank by closing the flow control valve. The related REFUEL SHUT OFF light turns on confirming that the flow-control valve has been closed by the pilot valve.

Gravity refueling is available through the wing mounted gravity refuel adapter on the top of the wing.

When DEFUEL is selected on the REFUELIDEFUEL master switch, the shutoff valves operate automatically to limit defueling to a preset quantity. If AC (variable frequency) power is available, the related auxiliary fuel pumps activate to aid the defueling process. Without an AC power source, suction defueling must be used. DC power is required for defueling regardless of AC power availability.

The long-range auxiliary fuel system is comprised of two auxiliary tanks, a refuel/defuel system, inter-tank fuel transfer system, venting system and an auxiliary tank control panel. Each auxiliary tank is located in the wing between the main tank and the fuselage. The long-range auxiliary fuel system increases capacity by 4,566 lbs (2,070 kg) of useable fuel, to a total capacity of 10,244 Ibs (4,644 kg).

Fuel quantity indication for the auxiliary tanks is provided in the flight compartment, by two MASTER FUEL QUANTITY indicators installed in the lower right hand comer of the COPILOTS INSTRUMENT panel. There are also two repeater indicators on the REFUEL/DEFUEL CONTROL panel. A magnastick located in each auxiliary tank provides an alternate means of measuring quantity on the ground.

Both auxiliary tanks can be refueled from over the wing independently or through the pressure refuel/defuel system. Fuel is transferred from each auxiliary tank, to its respective main tank by means of the inter-tank transfer system. Each auxiliary tank is vented overboard through the surge bay and two NACA vents on the underside of the wing.

The auxiliary tank control panel includes two MASTER FUEL QUANTITY indicators, two FUEL TRANSFER switches labeled transfer to mains to TANK 1 or to TANK 2 and OFF, a QUANTITY TEST push button and advisory lights.

Fuel transfer on both sides when applicable should be started immediately after takeoff, by selection of the transfer switches to TANK 1 and to TANK 2. This selection turns on the auxiliary boost pumps and opens a transfer valve in each main tank, which allows motive flow through an ejector pump to draw fuel from the auxiliary tank to the main tank. The TRANSFER TO MAINS switches must be turned off when transfer is completed.

The Dash 8 Q300 has three hydraulic systems, two main systems, and a alternate landing gear extension system.
The No. 1 and No.2 main systems power the:
• Flaps
• Landing Gear
• Nosewheel Steering
• Spoilers
• Rudder

The alternate landing gear extension hydraulic system is used as a backup for main landing gear extension.

Hydraulic power is provided by two independent systems, designated No.1 (left) and No.2 (right). The No.1 and No.2 hydraulic systems are normally pressurized by a related Engine Driven Pump (EDP). A Standby Hydraulic Pump provides backup power for the No.1 and No.2 systems. A Power Transfer Unit (PTU) and a manually operated hand pump provide hydraulic power for emergency purposes. An alternate rudder power system (ARPS) provides rudder control ifthe No.2 EDP fails or both systems lose hydraulic fluid.

The No. I 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:
• Flaps
• Normal Brakes
• Inboard Roll Spoilers
• Rudder (Lower PCU)
• Power Transfer Unit (PTU)

The No.2 system powers the:
• Landing Gear operation
• Outboard Roll Spoilers
• Nose Wheel Steering
• Emergency/Parking Brakes
• Rudder (Upper PCU)

The hydraulic system reservoirs store hydraulic fluid and supply the necessary fluid volume to the hydraulic systems. The #1 hydraulic system reservoir has a capacity of 2.68 U.S. Quarts (2.53 litres), and is installed in the No. I nacelle. The No.2 reservoir has a capacity of 5.19 U.S. Quarts (4.90 litres), and is installed in the # 2 nacelle. The reservoir uses system output pressure to pressurize itself (bootstrap) and provide suction 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.

Individual hydraulic pressure and quantity indicators are provided to monitor each system. Each quantity gauge is slaved to a mechanical master in each main wheel well.

Pulling out the ENGINE 1 or ENGINE 2 PULL FUEL OFF handle, will shut off hydraulic fluid to the related EDP. Hydraulic fluid is still available to the subsystems through the standby hydraulic pump.

#1 and #2 standby hydraulic pumps are used to supplement the related #1 and #2 main systems during takeoff and landing, or if an engine driven pump fails. Nominal operating pressure for the standby hydraulic system is 2900 psi.

The terms Standby Power Units (SPU) and electric standby hydraulic pump are one and the same, and are used interchangeably.

When the STBY HYD PRESS switches are set to 1 and 2, the Stanby Hydraulic Pumps are energized.

The No. I Standby Hydraulic Pump automatically operates if:
• Flaps selected greater than 0°

The No.2 Standby Hydraulic Pump automatically operates if:
• Flaps selected greater than 0°. or
• Loss of No.2 hydraulic system pressure. and
• No.2 condition lever selected to FUEL OFF. and
• weight off wheels

During takeoff and landing, #1 and #2 standby hydraulic switches are selected to positions 1 and 2 (on). If an engine or an EDP fails, hydraulic pressure is maintained by the related standby hydraulic pump. The # 2 standby pump also provides hydraulic pressure for an Alternate Rudder Power System (ARPS)(Mod 8/1983).

After takeoff. with FLAPS selector set to 0° and both switches set to NORM, the standby pumps stop.

During flap selections with the #1 EDP not operating, #1 hydraulic system pressure will fluctuate. This is due to the low volume output of the standby pump. This may cause the FLAP POWER caution light to come on, intermittently.

If either Standby Hydraulic Pump overheats. the related #1 or#2 STB HYD PUMP HOT caution light comes on.

Electrical power for #1 standby pump is supplied by the right 115 VAC bus. Normally electrical power for the #2 standby hydraulic pump is supplied by the left AC bus and monitored by its logic relay system. If left AC bus fails, relay logic switches electric power to the right AC bus. This causes the #2 SPU AUX PWR caution light to come on.

The primary purpose for the PTU, is to assist in raising the landing gear if the No.2 engine fails at takeoff. The PTU uses hydraulic pressure from the #1 system to power a hydraulic motor that then operates a hydraulic pump in #2 system. Hydraulic fluid is not shared or transferred between hydraulic systems. Hydraulic fluid must be available in the No.2 system for PTU operation.

The PTU may be selected on manually, or automatically. Manual selection is by pushing the MANUAL PTU SEL switchlight. The green light comes on when the PTU is outputting pressure.

The PTU automatically operates if:
• No.2 engine oil pressure decreases below 40 psi. and
• landing gear in transit up

PTU automatic operation stops when the landing gear is fully retracted.

Manual PTU selection is required to lower the landing gear when No.2 engine is shut down or the No.2 EDP fails.

CAUTION as per checklist
Do not select MANUAL PTU on if loss of No.2 system fluid has occurred.

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 selector valve, and gives access to the hand pump socket. Operating the hand pump lever, provides down pressure to the alternate landing gear actuators. This is used if the gear did not freefall into position during an alternate extension.

The alternate landing gear extension system is operated by a hand pump socket and lever. 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 as per checklist. 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 alternate rudder power system (ARPS) provides hydraulic power to the No.2 rudder system if No. 1 and No.2 hydraulic system fluid is lost, provided at least one AC generator is still operating. Two low pressure switches monitor system pressure and either one activates the No.2 standby hydraulic pump in order to power the rudder.

During flight, if the No. 2 EDP fails, the No. 2 standby hydraulic pump activates automatically and pressurizes the No. 2 hydraulic system. However, if the demand on the No.2 system is high, system pressure may decrease enough to cause the rudder shutoff valve to close. To prevent this, the PTU is manually selected on during approach and landing when the No. 2 engine is shut down (See Volume 1, Section 2). This supplements No. 2 standby pump pressure.

If No.2 system loses its fluid, the rudder shutoff valve closes and activates the rudder isolation hydraulic circuit to maintain rudder control.

Minimum dispatch # 1 system: 1.5 US Quarts Minimum dispatch #2 system: 3.0 US Quarts

Some aircraft have been modified with MOD 8/2781 as an alternate to MODS/1983 for protection of the hydraulic rudder system.

The hydraulic rudder isolation system (HRIS) provides hydraulic power to either the No. 1 or No.2 rudder actuators when both hydraulic systems fail and at least one engine is operational. The HRIS is engaged automatically whenever the fluid volume in its respective reservoir is below a minimum quantity.

A reservoir level switch senses reservoir volume when it is below the minimum dispatch level. When the reservoir piston reaches a preset level. the switch actuates. energizing the Rudder Isolation Valve (RIV) which subsequently closes both the pressure and return lines simultaneously. Closure of the valve prevents hydraulic fluid from flowing into the system containing all hydraulic services except the rudder. When a rudder isolation valve closes, it will turn on either the # 1 or # 2 HYD ISO VLV caution light.

From the point of actuation of the HRIS, a minimum of one hour flying time is available before the HRIS is disengaged. Disengagement occurs automatically when the reservoir is nearly empty.

The actuation point of the HRIS was selected below the minimum dispatch quantity of the smaller reservoir (System # I) so as to preclude the possibility of inadvertent actuation during normal operation.

If the #2 isolation valve close inadvertently. it will be necessary to extend the gear using the alternate system.

The Dash 8 Series Q300 Aircraft is approved for flight into known icing conditions. Ice and rain protection includes deicing, anti-icing, and rain removal systems.

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 lip.

The anti-icing systems use electrical heating elements to prevent ice formation. The system heats:
• leading edges of the propeller blades
• left and right pitot probes
• left and right static ports
• left and right stall warning AOA vanes
• left and right engine intake flanges
• both windshields and pilot's side window
• elevator horns

Electrically operated windshield wipers remove rain from the windshields.

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 taken from each engine and is available to inflate the boots regardless of the position of BLEED switches. System pressure is regulated to 18 psi and shown on the DEICE PRESS indicator, on the copilot's side panel. An isolation valve interconnects the two systems and is controlled by BOOT AIR switch which is normally open to ensure uninterrupted operation of either system, regardless of engine failure. The ISO position can be used to check regulated pressure in each system individually or to isolate a system leak.

Regulated deicer pressure is also used to inflate the passenger door seal, and to operate ejectors for the pressurization system outflow valve and the engine drains.

The boots inflate, 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 AUTO selector controls automatic boot operation, when selected to SLOW (4 min) or FAST (1 min). The operation is self-homing, such that a selection to SLOW or FAST and back to OFF will complete a full cycle. The boot inflation sequence begins and provides a dwell period appropriate to the selected rate. Green WING, TAIL and nacelle inlet lip, lights indicate boot inflation, sequence and confirm correct boot inflation pressure.

Integral distributor valve heaters are manually controlled by the VALVE HEAT switch, to prevent freezing.

Monitor ice accumulation between boot cycles (using wing inspection lights as required) to confirm that the selected AIRFRAME AUTO rate [FAST or SLOW] is appropriate.

If a malfunction occurs in the automatic timer, or a leak occurs in the system, the boots can be cycled manually with the AIRFRAME MANUAL selector. Rotating the selector through each of the six detented positions duplicates the automatic inflation sequence. Each set of boots (Wing and Tail) will inflate as long as the selector stays at the selected position. The related WING and TAIL boot inflation lights indicate full inflation.

Each selection should be held until the related pair of lights come on before moving to the next position.

If a leak occurs to the pneumatic lines of either deice system, the affected side will have to be isolated to make sure that deice pressure is available to the operating side. If the BOOT AIR switch is set to ISO, the isolation shutoff valve closes, isolating the failed side from the operating pressure side.

The engine intake boot on the side with normal pressure can be deiced.

All the boots on the horizontal and vertical stabilizers are pneumatically cross connected through a restrictor.(Dual sourced) However, the intake boot on the failed side is not.

When DEICE pressure to one wing is isolated following a rupture, the system must not be operated in AUTO or the wings will be deiced asymmetrically. To minimize lateral control problems while keeping the unaffected engine intake and tail clear of ice, the system should be operated in MANUAL with selection limited to the engine intakes and the tail boots.

The DEICE PRESS caution light will come on if the main deice pressure on either side is less than 5.5 psi.

The propeller blade leading edges are protected from ice accumulation by electrically heated elements bonded onto each blade. Electrical power is supplied from related left and right AC buses.

Propeller deice sequencing is controlled by one of two PROP TMR selector positions. Timer # 1 is normally used with #2 as a backup. Each timer can be selected to cycle at one of two rates, depending on the observed ambient air temperature. Activating either timer energizes the heaters of two blades of each propeller in a timed sequence as per Table XX. Four PROPS deice indicator lights monitor the sequencing and confirm blade heating.

TIME TIME SEQUENCE No. 115V PHASE ABOVE -10°C BELOW -10°C  on 10 or on 20


To protect the propeller heaters from possible overheat, the Proximity Switch Electronics Unit (PSEU) prevents heater operation when the aircraft is on the ground

The effectiveness of the propeller deicing system can be improved and propeller vibration reduced by operation of the propellers up to 1200 RPM.

On pre-Mod 8/0202 airplanes propeller RPM should be kept above 1000 RPM during propeller deicing operation.
On pre-Mod 8/0312 airplanes the SYNCHROPHASE switch must be set to OFF when propeller deicing system is operating.

Each pair of blade heaters can be tested on the ground by holding the PROP TEST switch to TEST, while the PROP TMR selector is turned to either #1 or #2 timer, (the test switch bypasses the PSEU). The four green PROPS indicator lights will come on in sequence to confirm heater/timer operation.

During ground test, do not hold PROP TEST switch at TEST for more than one cycle without a 60 second dwell time between cycles.

There are fuselage ice protection panels, on both sides of the fuselage, in line with the propeller arc. These prevent fuselage damage, in case ice is thrown off from the propellers.

An electric heater is installed in the intake flange of each engine. The heaters are powered by AC. Heater operation is confirmed by the HTR segment (amber), on the ENGINE INTAKE switchlight .

The heaters are energized when:
• the engine intake bypass doors are opened
• oil pressure in related engine
• temperature is less than + 15°C
• AC power is available

The aft part of engine intake flange is also heated by engine oil. Oil temperature must be kept above 45°C to make sure there is air inlet ice protection. The engine oil temperature gage should be monitored in icing conditions.

The left and right pitot heads and the left and right static ports have integral heaters to prevent ice build up. PITOT STATIC 1 is powered from the left essential bus. PITOT STATIC 2 is powered from the right main bus. They are controlled by the PITOT STATIC 1 and 2 switches, on the ICE PROTECTION panel.

The related PITOT HEAT 1. or 2 caution lights come on if
• PITOT STATIC switch is set to OFF
• pitot head heater power source is not available

The left and right stall warning AOA vanes are heated to prevent ice build up whenever variable AC power is available.

If a stall warning/stick pusher computer fails, or electrical power for the vane heater fails, the related #1 or #2 STALL SYST FAIL, and the PUSHER SYST FAIL caution lights will come on. The AOA control indicator, on the ADI / EADI, will also be removed from view.


The left and right windshields and the pilot's side window are heated to supply anti-icing and demisting. When the WINDSHIELD HEAT selector is turned to WARM UP, both windshields are heated with limited AC power. When the selector is turned to NORM, each windshield is heated with full AC power from its related AC bus. The left side window is heated when the PLT WOO/HT switch is set to ON.

Windshield temperature is regulated by individual controllers. An overheat condition shuts off power to the related windshield or side window heater. If a windshield overheats, the related L WSHLD HOT or R WSHLD HOT caution light comes on. If the pilot's side window overheats, the SIDE WOO HOT caution light comes on.

The windshield heaters stay deactivated until the caution light( s) go out and the HEAT selector is turned to WARM UP. When the side window cools to the normal operating temperature, the SIDE WOO HOT caution light goes out, and the window starts to receive heat again.


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.

An ice detector spigot is installed on each windshield wiper arm for inspecting possible ice accumulation. Momentary white pushbuttons, one on each side console, are pushed to light the related spigot in dark conditions.

The elevator horn leading edges have heater elements to prevent ice build up in the gap between the horns and tips of the horizontal stabilizer. Each element is heated by AC power. Setting the ELEV HORN switch to HEAT arms the control and warning circuits to activate the heaters. The heaters receive continuous voltage for heating, and use airflow cooling to prevent overheating.

To prevent possible overheating PSEU prevents the heaters from being energized until the aircraft is airborne, or is above 15°C (60°F).

For ground testing, the PSEU and ambient temperature cut out function can be temporarily bypassed, by holding the ELEV HORN switch at the TEST position. This causes the L ELEV HORN HEAT and R ELEV HORN HEAT caution lights to come on. After a 2 second delay, the heaters are energized and the caution lights go off.

If either heater fails, the related L ELEV HORN HEAT or R ELEV HORN HEAT caution light comes on.

The Dash 8 Series Q300 has retractable dual wheel tricycle landing gear. Hydraulically operated nosewheel steering gives directional control during taxiing, takeoff and landing. Each main wheel has hydraulically powered anti-skid brakes. There is also an emergency/park brake system.

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.

The landing gear is powered by the #2 hydraulic system and controlled by the LANDING GEAR selector lever in the flight compartment. There is a Power Transfer Unit (PTU) to supply back-up hydraulic power to the #2 hydraulic system for landing gear operation.

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 Switch 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 multi-disc brakes with an anti-skid system. The nose gear retracts forward and has a steerable nosewheel. The landing gear is operated by the #2 hydraulic system and is controlled by the LANDING GEAR selector lever. There is an alternate (emergency) means of extension for the main landing gear and the nose gear. Advisory lights give extension/retraction and fail/safe information.

Each main gear has a pair of forward, center, and rear doors hinged to the nacelle structure. When the gear is up, the main wheels are fully enclose. With the gear down, the forward doors on each main gear stay open. The nose gear has a pair of forward and aft doors, which completely enclose the nose gear when the gear is up. When the nose gear is down, the aft nose doors stay open.

The Proximity Switch Electronic Unit (PSEU) monitors the actuation of the gear uplocks, down locks, gear doors, Weight On Wheels (WOW), POWER lever position, and flap selected position. All of the data supplied to the PSEU is used to control the functions of the gear.

The PSEU controls the landing gear, (non mechanical)gear doors and their advisory lights, by monitoring proximity and WOW sensors. WOW signals prevent gear retraction while on the ground. A WOW system malfunction, illuminates a WT ON WHEELS caution light. 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 horn 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 forward compartment of the airstair door. With the gear extended, the pins are inserted into the gear stabilizer struts.

Landing gear operation is controlled and monitored from the LANDING GEAR panel to the right of the ENGINE INSTRUMENT panel. The landing gear is selected UP or DN (down) by moving the LANDING GEAR selector lever. A selector lever lock release button 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 down lock 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 on the flight compartment floor.

Landing gear warning horns sound in the flight compartment if the HORN switch is held at TEST. They also sound if the gear is not down and locked and:
1)    • flaps >10"
    • Autofeather system OFF
    • POWER levers below take-off setting

2)    • both POWER levers areat or near FLT IDLE
• IAS < 130

3)    • one POWER lever at or near FLT IDLE
• other POWER lever advanced
• IAS < 130
• HORN switch not latched at MUTE

#3 is the only case in which the landing gear warning horns may be muted.

When the LANDING GEAR selector lever is selected UP, the hydraulic pressure from the #2 hydraulic system is applied to the retract side of the landing gear system through the landing gear selector valve. The hydraulic pressure opens the nose gear forward doors, retracts the nose gear and closes the nose gear rear doors which are mechanically linked to the nose landing gear. After nose gear retraction. the nose gear forward doors close hydraulically. At the same time as the nose landing gear retraction sequence is occurring, the hydraulic pressure opens the main landing gear rear doors, which cause the main gear centre doors to open because the two are mechanically linked. Then the main landing gear retracts with the main gear forward doors which are mechanically linked to the main landing gear. After the retraction of the main landing gear. the main gear rear doors are closed hydraulically which causes the main gear centre doors to close.

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 out to show the gear is not locked down. The amber DOOR advisory lights come on to show the hydraulically operated gear 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 gear DOOR advisory lights go out to show all the hydraulically operated doors have closed. The solenoid selector valve is depowered, thereby hydraulically isolating the landing gear system. The main and nose gear are held in the up position mechanically.

The PSEU, through various sensors, controls the landing gear selector valve and the door sequence valves which control the related gear doors. If the door selector valves are not in the correct position after gear retraction. the LDG GEAR INOP caution light will come on.

The landing gear selector lever must not be used to extend the landing gear through the normal procedures when the LOG GEAR INOP caution light has come on.

When the LANDING GEAR selector lever is selected DOWN (DN), the hydraulic pressure is applied to the extend side of the landing gear system through the landing gear selector valve. The hydraulic pressure opens the nose gear forward doors and the nose landing gear is extended with the nose gear rear doors that will remain open. The nose gear rear doors are mechanically linked with the nose landing gear. At the same time as the nose landing gear extension sequence is occurring. the hydraulic pressure opens the main gear rear and centre doors. The main gear centre doors are mechanically linked to the main gear rear doors. The hydraulic pressure extends the main landing gear and the main landing gear forward doors that are mechanically linked to the main landing gear. After the landing gear has been extended. it is then that the PSEU senses that the three gear are down and locked. Then the door sequence valves are reversed and the hydraulic pressure closes the nose landing gear forward doors and the main landing gcar centre and rear doors. The main landing gear forward doors remain open and the landing gear is down and locked.

The advisory light sequence during extension starts with the LEFT, NOSE, and RIGHT red lights and the amber gear selector handle light coming on. The amber DOOR advisory lights 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 advisory lights go out and the green lights come on. The selector handle light goes out. The gear DOOR advisory lights go out when the hydraulically operated doors are closed. The solenoid selector valve remains powered to allow for continued hydraulic pressure acting on the gear when down and locked, however. primary down locking is maintained through over-center locks.

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 # 2 hydraulic system pressure or fluid

The system has an INHIBIT switch, manual cable operated and gear door uplock release mechanisms, a bypass valve, and an independent alternate hydraulic system. This includes main gear auxiliary extension actuators, a reservoir, and a selector valve. The landing gear extension INHIBIT switch is located above the copilot, adjacent to the MAIN LANDING GEAR ALTERNATE RELEASE door.

The alternate gear extension speed is limited to 140 KIAS because all hydraulic gear doors stay open after the gear is fully down.

Setting the INHIBIT switch to INHIBIT disconnects powered from the gear selector valve, preventing normal gear down extention using the LANDING GEAR selector lever. Selecting the LANDING GEAR selector lever to down still allows correct down and locked indications, and antiskid operation test.

Pulling the LANDING GEAR ALTERNATE RELEASE door fully open, mechanically positions the bypass valve to bypass all normal extension actuators to return, and exposes the main gear release T-handle. Fully pulling the T-handle releases the main landing gear doors and uplocks. Then the main landing gear will free fall.

Opening the LANDING GEAR ALTERNATE EXTENSION door, mechanically closes the alternate landing gear selector valve, and gives access to the alternate extension hand pump socket and the nose gear release handle. If the main gear does not free fall into the down and locked position, it can be pumped down by inserting the extension handle, located behind the copilot's seat, into the pump handle socket. The main landing gear can then be fully extended and locked down by pumping hydraulic fluid from the alternate hydraulic reservoir, located in the right hand side of the nose compartment.

The LANDING GEAR ALTERNATE EXTENSION door, and the main LANDING GEAR ALTERNATE RELEASE door must remain fully open during and after the alternate gear extension. The nose gear release handle is forward of the main gear alternate extension hand pump. When the release handle is pulled, the nose gear uplock and doors are released and the nose gear free falls to a down and locked position.

Setting the downlock verification switch to ON activates the downlock verification system. Three green lights indicate the related gear is down and locked. These lights are activated by an infrared system which is totally independent of the regular landing gear advisory system.

If the fluid quantity in the #2 hydraulic system is normal, but the #2 Engine Driven Pump (EDP) is not functioning, as may occur after a #2 engine failure, the Power Transfer Unit (PTU) may be used to supplement the pressure available to lower the landing gear. For this procedure, the PTU must be turned ON by pushing the PTU SEL switchlight before lowering the landing gear. A light in the switch comes on indicating the PTU is producing pressure. The PTU should not be used at any time during or following an alternate landing gear extention.

Directional control on the ground is by the Nosewheel Steering (NWS) system, powered by the #2 hydraulic system. Steering control is by either the steering control handle (tiller) or the rudder pedals. The tiller turns the nosewheel up to 60° either side of center for low speed taxi. Steering with the rudder pedals turns the nosewheel up to 7° either side of center for high-speed taxi, such as during takeoff and landing roll. After takeoff, 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 60° 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 Electronic Control Unit (ECU) is powered, when the nose gear is down and locked with weight-on-wheels.

Nosewheel steering using the hand control is limited to forward taxiing only.

With the nosewheel STEERING switch set to STEERING, the ECU is powered, if the nose gear is on the ground (weight-on-wheels).

The nosewheel will revert to a passive shimmy dampened castoring mode if:
• The nosewheel angle is more than 60°
• The ECU fails
• Steering is set to OFF

In the passive mode, the nosewheel will castor up to 120° 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
• ECU fails, or
• Nosewheel forced more than 60°

The NOSE STEERING caution light does not come on if electrical power is removed from the ECU.

To prevent the possibility of damage to steering actuators, reverse taxi is limited to a straight back motion with nosewheel STEERING switch in the STEERING position. Before taxiing in reverse, roll the aircraft forward to center the nosewheeL

If taxiing in reverse, DO NOT USE:
• Rudder Pedal or Tiller Steering
• Asymmetric Thrust
• Differential Braking
• Brakes to stop the aircraft

Taxiing in reverse should only be conducted on paved surfaces and in crosswinds less than 10 knots.

Each main wheel is equipped with a multi-disc brake unit powered by the #1 hydraulic system. An anti-skid unit modulates the application of brake pressure to each brake unit.

Brake pressure is applied by pushing the pilot or copilot BRAKE pedal. The anti-skid control unit:
• 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 GLARE SHIELD panel, operates the anti-skid system when selected to the ON position and the wheel speed is more than 10 knots. A self-test of the anti-skid control circuits is made when the switch is moved from the OFF to the ON position, or when the switch is selected to the momentary TEST position. A self-test, with the aircraft on the ground, causes the INBD ANTISKID and the OUTBD ANTISKID caution lights to come on for six seconds and then go out. A self-test in the air, with the landing gear extended and locked, will turn the caution lights on for three seconds. Failure of the anti-skid control circuitry turns on the related caution light continuously.

The PSEU supplies weight-on-wheels and gear up and locked signals to the anti-skid control unit, to ensure that the brakes are off until the aircraft has touched down and the wheels are spinning. In conditions of low runway friction, the main gear wheels may not spin up before the aircraft's weight is fully on the wheels. In this case, the anti-skid control unit gives a 3 second delay before brake pressure is applied. This delay is immediately cancelled when wheel speed is more than 35 knots.

Aircraft operation with the anti-skid brake control system inoperative is not permitted when on unpaved surfaces. The ANTI SKID switch must be set to OFF. See Supplement 26.

Brake cooling times (Volume 3, Chapter 3, and Section 11) must be observed between a landing, or a low energy rejected take off and a subsequent take-off, to make sure that sufficient brake energy is available to bring the aircraft to a complete stop if the subsequent take off is rejected.

The emergency/parking brake system is used to set the parking brakes, and also, as an alternate means of stopping the aircraft when the normal braking system is not available. An EMERG BRAKE lever on the engine control quadrant operates the system. The emergency/parking brake system is powered by the #2 hydraulic system or by the hydraulic accumulator system. Hydraulic pressure to the emergency/parking brake system is indicated on the PARK BRAKE pressure indicator on the copilot's instrument panel. The gauge also indicates the brake accumulator pressure if #2 hydraulic system pressure is unavailable. A minimum accumulator pressure before engine start is 1500 psi. A hand pump, located in the right main wheel well, can be used to charge the accumulator.

The parking brake is engaged by pulling the EMERG BRAKE lever back to the detent, PARK position. This turns on the PARKING BRAKE caution light. To release the parking brake from the detent, push the button on the side of the EMERG BRAKE.

For aircraft braking, pull the EMERG BRAKE lever aft. The lever operates against a spring to produce a resistance that is proportional to the brake pressure applied. Brakes should be applied with a steady continuous pressure. If # 1 hydraulic system is not available and the #2 hydraulic system is not serviceable, the accumulator will provide the brake pressure. A fully charged accumulator allows for approximately four full applications. Differential braking and anti-skid protection are unavailable when using the emergency parking brake system.

Care should be taken when releasing the lever as considerable spring tension will force the lever forward abruptly.

With the parking brake set, application of engine power will cause the take-off warning horn to sound.

The flight compartment lights  include:
• Dome lights
• Storm lights
• Panel and Instrument lights
• Pilot's and Copilot's Panel lights
• Pilot's and Copilot's Utility lights
• Pilot's and Copilot's Map lights
• Circuit Breaker Panel lights
• Observer's Utility light

The DOME lights are directly connected to the battery bus.

To avoid loss of battery power, the dome lights switch must not be left in the DOME position unless a source of power is available.

When set to STORM/DOME, the storm and dome lights come on.

Flight compartment panel and instrument integral lighting is supplied by disc shaped lamp assemblies embedded in Plexiglass.
There are panel lights for the:
• Overhead Console
• Glareshield panel
• Engine Instrument Panel
• Center Console

The clock displays are dimmed by a switch on each clock. The standby compass light is controlled by the CAUT/ADVSY LIGHTS DIM BRT toggle switch.

Two swivel ball utility lights are dimmable by an adjacent knob, with an OFF position.

The pilot and copilot's map lights can be dimmed using a related knob adjacent to the map light.

two white floodlights.

on the flight compartment bulkhead.

• Airstair door lights
• Boarding lights
• Cabin overhead lights
• Cabin sidewall lights
• Buffet overhead light
• Galley lights
• Wardrobe overhead light
• Lavatory compartment lights
• Passenger reading lights

All PSUs may be tested with the PSU TEST switch on the attendant's panel

• No smoking
• Fasten seat belts
• Lavatory availability

A Return to Seat sign is located in the lavatory. Passenger cabin exit locators and exit signs are part of the emergency lighting system

They also come on automatically when the landing gear is selected down. A low tone chime sounds on the PA when the signs come on and a pictorial warning notifies passengers to not smoke.

With option 825CH01356 installed, the NO SMOKING signs m'e always on when battery power is available, even if the NO SMOKING switch is set to OFF. The switch will still sound a low tone chime when it is set to the ON position.

A RETURN TO SEAT sign in the lavatory compartment also comes on with the fasten seat belt signs. The FASTEN BELTS toggle switch is on the overhead console. A low tone chime sounds on the PA when the signs come on.



The baggage compartment has two lights, a door light adjacent to the top of the baggage door, and a light on the compartment ceiling

If the lighted BAGGAGE CaMP switch, on the attendant's control panel  is pushed, both lights come on. If the baggage door is unlocked, the baggage compartment lights also come on,

• Nose compartment service light
• Tail cone service light
• Electrical equipment rack service light
These lights come on when the access door is open and the rotary switch at the end of the light is tumed on.


The REFUELIDEFUEL panel light comes on when the panel door is opened.

The exterior lighting system is used for:
• Aircraft visibility
• Visual inspection of the aircraft exterior
• Lighting the runway for approach and landing

The aircraft wing tips and vertical stabilizer lights include:
• An anti-collision ramp light on top of the vertical stabilizer (Red)
• An anti-collision strobe light on each wing and on the tail cone (White)
• Primary and secondary wing position lights (Red, Green)
• An upper tail position light (White)

Wing and engine inspection lights provide visual inspection of their respective surfaces.

Other exterior lights include:
• Landing approach lights
• Landing flare lights
• Taxi light
• Optional logo lights

The dual red and green lights mounted beside each other in the transparent wingtips are referred to as primary and secondary lights. The primary light is mounted in the forward position. When the POSN LIGHT switch on the right EXTERIOR LIGHTS panel, is first set to POSN, all position lights come on. Approximately two seconds later, the secondary lights go out but remain anned, and come on if the related primary light fails.


The flashing red anti-collision ramp light is located on top of the vertical stabilizer

There are two white flashing anti-collision strobe lights in each transparent wingtip. and two on the fuselage tail cone
LOGO LIGHTS optional

The emergency lights supply interior and exterior lighting for use in emergency situations, or if the 28 VDC R SECONDARY bus power fails. consists of
• Two switches
• Emergency lighting control box
• Portable emergency light and power supply
• Passenger address (PA) system switching relay
• Six separate emergency lights power supply units (each with its related group of lights)

The emergency lights are controlled by the flight compartment EMER LIGHTS three position (OFF, ARM, ON) switch, and the attendant's EMERGENCY LIGHTING pushbutton, adjacent to the attendant's seat. The attendant's EMERGENCY LIGHTING switch is an alternate action pushbutton, and has two operating positions marked NORM (green) and ON (yellow). When armed, the emergency lighting activates if the right secondary DC power is interrupted. The PA system power is also switched to the battery bus when the emergency lights are on. locations include:
• Emergency exit egress lights
• Airstair door egress light
• Emergency exit signs with threshold lights
• Escape path lighting strip
• Mid-cabin exit locator with downlights
• Ceiling aisle lights
• Flight compartment emergency light
• Forward cabin exit locator with downlights

The six emergency lights power supply units are mounted in the cabin ceiling and in the wardrobe. The batteries in the emergency lights power supply are charged from 28 VDC power when the EMER LIGHTS switch is in either ARM or OFF, and the attendant's EMERGENCY LIGHTS switch is in NORM.. Fully charged batteries power the emergency lights for approximately 20 minutes.

A floor proximity emergency escape path system is installed as a visual aid in case of emergency evacuation. The emergency escape path system consists of lighted arrows placed adjacent to each seat from the rear of the passenger cabin to the middle exits. Exit locations are designated by a series of four closely spaced red lights with red lighted EXIT signs at the exit doors.

The portable emergency light on the flight compartment ceiling, aft of the OVERHEAD console. The portable light is removed by pulling the release handle labeled PULL EMERGENCY LIGHT,. The batteries are on continuous charge when the airplane's 28-VDC power is on, the EMER LIGHTS switch is in the ARM position, and the attendant's EMERGENCY LIGHTS switch is in the NORM position. When a charging current flows through the batteries, the dual-lamp indicator comes on, showing a faint light through the lens. A test button, when pushed, verifies emergency light operation

The flight compartment EMERGENCY LIGHTS switch must be set to OFF, and the attendant's EMERGENCY LIGHTS switch to NORM, to disarm all the emergency lights power supply units and the portable emergency light and power supply before switching off airplane electrical power. This prevents power failure circuits in the units £i'om energizing the lights and discharging the batteries.

There are fixed and portable oxygen systems for the flight crew. A separate system is provided for the passengers. Optional observer's fixed oxygen and first aid oxygen are also available.

The crew fixed oxygen system includes two or three microphone equipped pilot masks with a controlled oxygen dilution regulator, Portable crew, passenger and optional first aid oxygen systems have regulators with ONOFF selectors. Optional Protective Breathing Equipment (PBE) units m'e available for the flight crew and passengers.

The fixed crew oxygen system  provides a source of supplemental oxygen sufficient to permit descent from 25,000 feet to 14,000 feet in four minutes and continued flight at 14,000 feet for 116 minutes, Oxygen can be delivered as either 100% or diluted with ambient air. The quick-donning type oxygen masks include a microphone and regulator. The masks are suspended from quick release straps on the bulkhead behind the pilot and copilot seats. Each mask plugs into its associated oxygen outlet and each is supplied from a single, common cylinder in the right hand lower nose. When installed, the optional observer's mask is supplied from the copilot's oxygen supply line.

System oxygen pressure is shown on a gauge located on the copilot's side panel A green over-pressurization disc is located on the right side exterior of the nose. The disc is ejected out if over-pressurization occurs in the cylinder.


If the mask is donned and difficulties in breathing are experienced or the oxygen line indicator is RED, make sure the supply hose is connected.

On pre-Huntington interior airplanes, a portable oxygen cylinder with fullface smoke mask It is located on the bulkhead behind the copilot's seat. The cylinder can supply a flight crew member with 15 minutes of oxygen at a pressure altitude of 8,000 feet.

Post Huntington interior airplane operators will have Protective Breathing Equipment (PBE) installed in place of the crew portable oxygen system.

An optional attendant oxygen cylinder is available. It is kept in the the right forward overhead bin. When the cylinder is fully charge to 1800 psi, it can provide 30 minutes of oxygen.


The passenger portable oxygen system has three dual outlet oxygen cylinders and six masks The cylinders and masks are stowed in the emergency compal1ment at the rear of the buffet.

Each cylinder contains:
• ON-OFF selector
• Pressure indicator
• Safety discharge outlet with li'angible disc
• Charging valve
• Two quick-disconnect outlets for passenger masks.
A carrying strap is attached to each cylinder. The masks are of the reservoir bag type.

Each cylinder can sustain two passengers for a minimum of 30-minutes. Each cylinder has an altitude compensated continuous flow regulator such that no oxygen is available below 10,000 feet cabin altitude.

With the cabin pressurized, this cylinder is unsuitable for first aid use

1300 psi (2 Crew) 1800 psi (3 Crew)
1050 psi (2 Crew) 1450 psi (3 Crew)

• Crew portable oxygen system
1750 psi

• Passenger portable oxygen system
1100 psi

• First Aid oxygen system
1100 psi

The pneumatic systems use bleed air from the engines or by an optional Auxiliary Power Unit (APU). The engines supply bleed air for pneumatics after engine start. If engine bleed air is not available, the APU can supply bleed air for air conditioning on ground.

The pneumatic system supplies air for:
• Air conditioning
• Aircraft pressurization
• Airframe deicing
• Door seal pressurization

APU bleed air is used for:
• Air conditioning

Bleed air operation from the left and right engines are similar. Compressed bleed air is ducted from the Low Pressure (LP) compressor port, or the High Pressure (HP) compressor port of each engine. POWER lever position and pressure switches in the system determine whether the HP port or the LP port will be used.

Setting the two BLEED control switches, on the AIR CONDITIONING control panel, to 1 and 2, turns on the related engine bleed air system. A single rotary BLEED selector, labeled MIN/NORM/MAX helps controls bleed airflow to the air conditioning system.

Pushing the bleed air (BL AIR) switchlight, on the APU CONTROL panel, turns on the APU bleed air.

After engine start, the BLEED switches are set on and the BLEED air flow control knob is turned to the MAX position. The two engine bleeds join downstream of the wing isolation check valves and provide bleed air to an air conditioning pack in the aft fuselage. It is pressure regulated by a pressure regulator.

At high power settings, (during takeoff and cruise), bleed air comes from the LP port. At low power settings, bleed air comes from the HP port.

The two BLEED switches are powered from the left and right main distribution buses. Each switch, when set to 1 and 2 will open the Nacelle Shutoff Valve (NSOV) and pressure regulator valve.

Bleed air is continuously available to the deicing system, regardless of the position of the BLEED switches.

The bleed flow control on the AIR CONDITIONING panel is a rheostat labeled BLEED with MIN and MAX. This rotary knob is powered from both the left and right secondary bus. The flow control system uses this duplicate supply to make sure proper flow selections are maintained if an engine fails .

The selection on the BLEED rotary knob sends an identical flow control signal to each of the servos (one mounted in each nacelle) which then commands a specific supply from the Pressure Regulator. Rotating the BLEED knob will vary the bleed flow volume.

The bleed air system in each nacelle consists of a precooler, two high pressure switches, a high pressure shutoff valve, high pressure venturi, low pressure check valve, nacelle shutoff valve, duct overpressure and overtemperature switches, and flow control servos.

Two high-pressure switches are connected duct from the engine high pressure bleed port. Each switch is a normally closed, pneumatic pressure sensitive switchs are connected into the circuit to the high pressure shutoff valve when the BLEED switch is set to BLEED. The No.1 switch is set at 65 psi; the No. 2 switch is set at 55 psi.

A venturi, installed in the high pressure bleed port, restricts the bleed air flow to a maximum of 10%,in case the duct ruptures.

The high-pressure bleed shutoff valve is line mounted in the hot air duct from the high-pressure bleed port on the engine. It is a normally closed valve, used to control the flow of high-pressure bleed air from the engine.

The precooler, heat exchanger is mounted above the engine in the high pressure bleed air supply line. Its purpose is to cool the hot HP engine bleed air to an acceptable temperature for parts downstream.
The cooling effect is achieved by low pressure bleed air, through the Handling Bleed Off valve (HBOV), flowing across the precooler and out the exhaust on top of the engine nacelle.

The HBOV is used to prevent compressor surge and stall, and to reduce noise level during steady operation of the engine.

If one engine remains at idle power as power is increased on the other, the HP shutoff valves will close on both engines when the HP outlet duct pressure increases beyond the set point on the high power engine. As power is reduced on the higher power side, both HP shutoff valves will open, restoring bleed supply from the high pressure ports.

If there is an electrical failure, the high-pressure shutoff valves will move fully closed.

The nacelle shutoff valve will deenergize open ensuring a constant supply of low pressure bleed air to the system. The bleed flow control will operate at a maximum flow setting regardless of rotary knob position.

The related #1 BLEED HOT or #2 BLEED HOT caution light comes on if an overtemperature condition is detected. The control circuit automatically shuts down the bleed air system by closing the related:
• High pressure shutoff valve
• Nacelle shutoff valve

Low pressure bleed air is still supplied to the deicing system.

When the temperature drops below limits, the caution light will go off, and the bleed valves will operate normlly. If a BLEED HOT caution light continues to cycle on and off, turn off the related BLEED switch, and monitor aircraft pressurization and heating closely for the remainder of the flight. As per checklist

An overheat shutdown of one engine's bleed does not affect the continued normal operation of the opposite engine's bleed.

The Dash 8 Series Q300 is powered by two Pratt & Whitney PW123E turboprop engines (Models 311 and 315). Each engine drives a four bladed, constant speed, fully feathering and reversible Hamilton Standard propeller. Each engine develops 2140 Shaft Horse Power (SHP) under normal take off conditions. An automatic uptrim feature allows either engine to develop a time limited maximum take off power of 2380 SHP, if the opposite engine fails during take off.

Model 314 aircraft, incorporating CR872CH00006, CR872CH00007, and CR872CH00008, have PW123B engines. Each engine develops 2250 Shaft Horse Power (SHP) under normal take off conditions. Automatic uptrim feature allows either engine to develop a time limited maximum take off power of 2500 SHP, if the opposite engine fails during take off.

The engine has a low pressure (first stage) and a high pressure (second stage) centrifugal compressor, attached to separate single stage turbines. A two-stage power turbine drives a third shaft to turn the propeller through a reduction gearbox. The high-pressure compressor drives the accessory gearbox.

POWER and condition levers are used for related engine and propeller control. The Engine Electronic Control Unit (ECU) and Mechanical Fuel Control (MFC) change engine power with input from the POWER lever. The POWER levers change engine power in the forward range, and also propeller blade angle in the idle through reverse beta range. The condition levers set propeller RPM in the forward thrust range, provide manual propeller feathering, and fuel on/off control.

Air entering at the engine inlet is directed rearward and compressed. Two centrifugal compressors carry out compression for combustion and bleed extraction purposes. Air is first ducted to the low-pressure (NL compressor and then to the high pressure (NH) compressor where it undergoes a second stage of compression. The air then enters internal ducts, is mixed with fuel, discharged into the combustion chamber and ignited. Gases exiting the combustion section initially impact onto a single stage NH turbine. The turbine extracts energy from the hot gases, and drives a shaft directly connected to the NH compressor. A gear drive attached to this compressor drives the accessory gearbox mounted on the top section of the turbo-machinery. Mounted behind the NH turbine is a single stage NL turbine, which also extracts gas energy. It drives a shaft connected directly to the NL compressor. As the combustion gases continue to flow rearward they are directed towards the two-stage power turbine assembly. The power turbines turn as a single unit, extracting the majority of gas energy remaining to rotate a shaft connected to the reduction gearbox. The reduction gearbox, at the front of the engine, then transfers the power to the propeller. After leaving the power turbine, the gases are vented through to the exhaust pipe, above and behind the wheel well and along the top of the nacelle. The gases are then are discharged overboard.

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

Each engine nacelle intake incorporates a bypass door, which provides a means of preventing solids and precipitation from entering the engine intake. Door operation is controlled by switchlights on the engine instrument panel
The doors are selected open during flight when any of the following conditions are encountered:
• Icing Condition
• Heavy Precipitation
• Bird Activity

On the ground they also are open when operating from contaminated runways.

Powerplant control is by POWER and condition levers, on the center console

The two POWER levers, marked 1 and 2, control engine power in the forward power range, and engine speed and propeller blade angle in the idle through reverse Beta range.

POWER Levers Select:
• Power for Flight
• Flight Idle (FLT IDLE)
• Reverse (MAX REV)

POWER lever movement from a point above FLT IDLE forward to MAX PWR selects higher fuel flows, giving increased engine power and speed.

As the POWER lever reaches a point slightly above FLT IDLE, it enters the flight Beta range.  In this Beta range they also control blade angle.

For descriptive reasons, the beta range of POWER lever movement above FLT IDLE is designated FLIGHT BETA (used while airborne), and the beta range between FLT IDLE and MAX REV is designated GROUND BETA (used while on the ground). A flight idle gate prevents unintentional movement of the POWER levers into the GROUND BETA region.

The gate is overridden by raising gate release triggers below the handgrips, allowing the POWER levers to be moved further aft.

A Beta warning horn (optional) will sound if the gate is raised in flight.

Further POWER lever movement aft moves the blades into reverse until the POWER levers reach MAX REV. Between DISC and MAX REV, fuel flow and power output is increased.


Two condition levers, marked 1 and 2 set:
• Maximum Propeller RPM (MAX)
• Propeller RPM
• Minimum Propeller RPM (MIN)
• Propeller Feather & Fuel On (START & FEATHER)
• Engine Shutdown (FUEL OFF)

For flight operation in the forward power range, a governor in the Propeller Control Unit (PCU), which regulates propeller speed (Np) controls propeller blade angle, in response to condition lever settings.

During take-off and in flight, in the MAX position, propeller RPM is governed at approximately 1200 RPM, and in the MIN position, propeller RPM is governed at approximately 900 RPM. Changing the Condition Lever Angle (CLA) between the MIN and MAX positions, changes the propeller speed between 900 and 1200 RPM. On the ground with POWER levers at FLT IDLE, and a CLA between MIN and MAX inclusive, the propeller speed is 785 RPM. This is known as prop underspeed governing. In the START & FEATHER position the propeller is feathered and GROUND IDLE operation of the engine is scheduled. Moving the lever to FUEL OFF stops all fuel flow to the engine. Lift gates prevent unintentional movement of the CL from MIN to START & FEATHER, and from START & FEATHER to FUEL OFF.

Protection from NH compressor rotor overspeed is provided at 105.8% NU

Engine operating information is transmitted to gauges mounted on the center engine instrument panel. The gauges provide indications in both analog and digital forn, and include the following:
• TRQ Engine Torque indicated as a percentage of the maximum.
• PROP Propeller speed indicated in RPM.
• NH NH turbine and compressor speed as a percentage of maximum.
• NL NL turbine and compressor speed as a percentage of maximum.
• ITT Inter- Turbine Temperature shovm in degrees Celsius. An amber light in the lower right corner of the gauge comes on when the temperature is 800DC or higher .
• FF Fuel Flow to the engine combustion section is shown in hundreds of pounds or kilograms per hour.

Each of the above instruments has a PUSH- TO-TEST button in the lower left corner. When the PUSH- TO- TEST button is pushed, the analog and digital scale show 1050 or 105.0, as applicable.

Dual split analog oil temperature and pressure gauges show engine oil pressure and temperature. The oil temperature display is in degrees Celsius and oil pressure in psi.

Each engine has an ignition system with two exciter units and two igniter plugs in the combustion chamber. The system is activated and deactivated automatically by the engine start control circuits during the start sequence.

Setting the ignition switches to MANUAL gives continuous ignition for flameout protection, for flight into precipitation, icing or turbulence.

Aircraft with ModSum 8QI00813 have an AUTO position in place of the NORM. The IGNITION switches are labled OFF-AUTO-MANUAL. The engine is started with the related IGNITION switch in the AUTO position. The AUTO position also arms an automatic relight system. If Nl drops below 60%, the igniters automatically operate. The igniters will continue to operate until Nil goes above 60'%. The AUTO ignition does not operate if the condition lever is in the FUEL off position.

The start system is armed by selecting the engine to be started on the ENGINE START panel and turned on using the START/ SELECT switch. The start sequence is initiated by pushing the START switchlight (START segment of switch light on). Engine starting is accomplished using the starter/generator in conjunction with the ignition and fuel control systems. The starter/generator rotates the High Press (NH) compressor through the accessory gearbox, to develop the necessary airflow and engine RPM before fuel is introduced.

As the starter/generator turns the high-pressure compressor rotor through 10%-19% NH, the condition lever is selected to the START & FEATHER position to introduce fuel, and engine starting occurs. The start sequence will automatically terminate at 63 ± 2% NH at which time the ignition system is deactivated, and the starter/generator is switched to the generator mode. As the stal1 circuit is disarmed the select switch is automatically released to the center position and the START segment of the start switchlight goes out. The engine continues to accelerate to ground idle, stabilizing at 72 - 78% NH.

The Engine Electronic Control Unit (ECU) and Mechanical Fuel Control (MFC) normally manage engine fuel control. The ECU calculates the most efficient fuel flow for the conditions and sends electric fuel control commands to the MFC. The MFC is mounted on the engine driven fuel pump and is responsible for physical metering of the fuel to the combustion section. The ECU electronically schedules fuel when the ECU mode switchlights, on the left side of the fuel control panel, are pushed ON (green).

The ECU receives both engine and air data computer inputs. It computes and compares data to set points in its memory and adjusts a motor in the MFC. This provides precisely metered fuel, to adjust engine power requirements for various operating conditions.

The ECU is used for:
• Ground Idle Fuel Schedule
• Prop Underspeed Governing 785 RPM (Beta Operation)
• Prop Governing - Reverse 1140 RPM
• Quiet Taxi Mode
• Fuel Scheduling Forward Power Range
• Power Uptrim Maximum - 2380 SHP
• Handling Bleed Valve Operation

During engine start, the ECU helps accelerate the engine to 72-78% NH

During Beta operation, the ECU schedules fuel to maintain prop speed at approximately 785 RPM. This ensures the propeller speed does not drop below 780 rpm (the restricted range). The underspeed governing can be cancelled if the condition lever is moved below MIN to START & FEATHER or FUEL OFF, or if the ECU is selected to MANUAL.

Aft of the DISC position the ECU increases fuel flow allowing propeller rpm to increase until the ECU is regulating prop speed at 1140 rpm. This occurs at the MAX REV POWER lever position.

A QUIET TAXI mode can be entered while the engine is in the beta range, to reduce propeller noise. By moving the condition levers to the MIN position, with the POWER levers at FLT IDLE, the ECU limits fuel flow to that required to keep propeller speed at the 785 RPM (Np underspeed governing limit), throughout the reverse POWER lever range. QUIET TAXI is cancelled when the condition levers are moved to MAX.

In the forward power range the POWER lever position tells the ECU to schedule fuel to achieve and maintain torque at a value equal to the POWER lever position.

Each ECU features a power uptrim mode. If an engine fails during take off, with autofeather selected, power output of the operating engine increases 10% of set power (to a maximum rating of 2380 SHP). A PWR UPTRIM advisory light on the engine instrument panel, will come on to show that UPTRIM is activated. Deactivation of the autofeather system after take off removes the UPTRIM signal. The PWR UPTRIM advisory light goes out and the ECU returns the engine to the original power setting.

The ECU automatically controls handling bleed valve operation in order to prevent compressor stalling and surging.

If the ECU fails, the fuel system goes into reversion (MANUAL) mode. Engine fuel control is now via the MFC only. The related # 1 ENG MANUAL or # 2 ENG MANUAL caution light comes on. The affected engine will increase up to 10% torque of the last set power. The POWER lever of the affected engine will have to be adjusted to restore power.

Full engine power is available when in manual mode, operation is characterized by increased engine responsiveness to POWER lever movement. If one engine is operating in the normal ECU fuel schedule mode, asymmetric POWER lever positions will result when setting torque values equally on the gauges.

The manual operating mode provides fuel scheduling for engine operation, but docs not supply corrected control functions associated with ECU control.

When operating an engine in the manual mode on landing, DO NOT retard the POWER lever below the DISC detent as per checklist. After landing, maintain propeller speed (Np) above 785 RPM and high-pressure turbine speed NH above 66%. If unable to observe these limits, feather the propeller after landing is complete and taxi aircraft on remaining engine.

The engine oil system provides lubrication of the engine bearings and gearboxes. and supplies oil for propeller operation . The integral oil tank in the engine (6 US gallons total capacity) has a sight gauge to check oil quantity. Oil temperature is controlled by the oil cooler that automatically opens and closes an oil cooler door (aft of the bypass door) based on oil temperature. When oil pressure drops below 44 psi. the related #1 ENG OIL PRESS or #2 ENG OIL PRESS warning light flashes.

The constant speed four-blade propeller, is fully feathering, and reversible. Bonded to the inner leading edge is an electrical heater element. The propeller is driven through a reduction gearbox, which also drives the:
• Propeller Control Unit (PCU)
• Propeller Control Unit Oil Pump
• Propeller Overspeed Governor
• Main Hydraulic Pump
• AC Variable Frequency Generator

A pitch change mechanism in the hub, receives increased engine oil pressure, supplied through a Propeller Control Unit (PCU), to control blade angle and therefore propeller speed. The propeller system provides:
• Governed Constant Speed Operation
• Power Lever Controlled Beta Range (Flight Idle to Reverse)
• Manual Feather
• Alternate and Unfeather
• Propeller Synchrophasing.

A propeller-restricted range exists between 500 and 780 RPM, within which continued operation of the propeller is prohibited due to vibration fatigue. During Beta mode operation. the ECU Np underspeed governing circuits prevent Np from decreasing below 780 unless canceled by propeller feathering. All propeller blade angles are achieved by controlled metering of oil pressure to the propeller actuator piston in the hub. Aerodynamic loads on the propeller tend to drive the blades to low blade angles (Higher Prop RPM). If the supply of oil pressure to the PCU be lost, a pitch-lock feature will prevent aerodynamic/centrifugal  loads from reducing blade angle more than 1 degree. and propeller RPM from increasing more than 2% from the point of oil pressure loss.

The propeller provides for governed constant speed operation through a propeller governor controlled by the condition levers. The POWER levers control blade angle in the beta range. The feathering is controlled by the condition levers or by the autofeather/alternate feather system.

When the POWER levers are in the forward power range, condition lever inputs to the propeller governor set the propeller rpm. The governor then controls propeller rpm by adjusting blade angle as required to keep the rpm constant. The propeller blade angle is adjusted by metering prop pump oil to one side of the propeller actuator piston, which is opposed by constant oil pressure supplied to the opposite side of the piston.

The Beta range is from a POWER lever position above flight idle (called FLT BETA) to the full reverse (MAX REV) position. When the propeller is in the Beta mode, blade angle is set by the POWER lever input. PROPELLER GROUND RANGE 1 and 2 advisory lights, on the left glare shield panel, indicate when the propellers are in the ground Beta range.

A beta back-up system
If a propeller enters the ground Beta range with POWER levers above the ground Beta range setting, the Beta backup protection system causes the blades to be driven coarse. The system is then restored to its original condition. If the problem persists, the result is a continuous cycling in and out of the propeller GROUND BETA range, until the fault is corrected, or the POWER levers are advanced to a position where normal governing is reinstated. The related PROPELLER GROUND RANGE light will flash as the propeller cycles.

The propeller overspeed governor has a hydraulic section and a pneumatic section.

The hydraulic section controls blade angles hydraulically when prop rpm is more than 1250 rpm. During an overspeed condition, propeller rpm is reduced by increasing blade angles. When propeller rpm decreases below the overspeed limit, the overspeed governor restores normal propeller governor control. If the propeller goes back to an overspeed condition, the cycle is repeated, resulting in a continuous fluctuation in prop rpm in and out of overspeed until the cause is removed.

The pneumatic section of the overspeed governor causes the MFC to reduce the amount of fuel being supplied to the engine, if an overspeed of approximately 1308 rpm is reached. Reducing fuel to the engine causes NPT to drop, in turn lowering the propeller rpm. When Np drops below the overspeed limit, the governor allows the MFC to restore normal fuel flow. If the propeller goes back to overspeed, the cycle is repeated, resulting in a continuous prop RPM and fuel flow fluctuation until the fault is removed.

The hydraulic section of the overspeed governor does not operate in reverse, therefore, the pneumatic section is the only propeller overspeed protection in this range.

The propeller synchrophase system synchronizes the RPM of both propellers in climb, cruise or descent. Synchrophasing is achieved by matching the RPM and phase of No. 2 propeller (SLAVE) to that of No. 1 propeller (MASTER). A SYNCHROPHASE switch, on the right glareshield, activates the system. Before selecting the system ON, the the propellers must be manually set to within ± 20 RPM of each other. A green advisory light, adjacent to the synchrophase switch, indicates the system is selected on.


The Beta Lockout System is in place to prevent the propellers from entering the ground beta range during flight.

During flight, if the PLA is placed below Flight Idle, the Beta Lockout System is activated. It is deactivated when the
RAD ALT less than 50' AGL,
the WOW shows the aircraft on the ground,
power levers are returned to flight idle.
A warning horn provides an aural warning as the Flight Idle gate is removed by lifting the power lever trigger, and before ground beta is actually selected.

With PLA less than Flight Idle, as propeller pitch decreases, Np will increase. If Np exceeds 1000 RPM. a DC signal to the feather solenoid will decrease the RPM. Once Np drops below 1000 RPM, the signal is removed. As long as the PLA remains below flight idle, the Np will fluctuate between 900 and 1100 RPM, preventing the pitch from decreasing into ground beta.

The Beta Lockout Test and the Beta Backup Test are included in the 24 hour systems checks to be completed prior to the first tlight of the day. Refer to your AFM for required test(s) based on aircraft modification status.

Propeller feathering systems provide:
• Autofeathering
• Manual feathering
• Alternate feathering andunfeathering.

automatically feathers the propeller if an engine fails during takeoff .
increases the power (uptrim) of the operating engine.

Selected on by pushing the AUTOFEATHER switchlight, on the engine instrument panel. This causes the SELECT segment to turn green. The ARM segment will come on when both engine torques are more than 35%, and both POWER levers are advanced to meet PLA/OAT requirements. If the arm light does not come on, the take off must be rejected.

The autofeather system is activated if the torque of one engine is less than 25%, while the torque of the other engine is more than 35%, and both POWER levers are advanced to meet PLA/OAT requirements. Uptrim is immediately initiated on the good engine through that engine's ECU, and the PWR UPTRIM advisory light comes on.

After a 3-second delay, autofeathering of the failed engine begins by
energizing the feather solenoid valve,
canceling the Np underspeed signal to the ECU, and
activating an auxiliary feathering pump for approximately an 15-second period to sure adequate oil pressure is available for propeller feathering.
The auxiliary (feather and unfeather) pump provides a backup source of oil pressure to the propeller pitch-change mechanism. The pump uses oil from an auxiliary oil reservoir in the propeller RGB, to permit autofeather if engine oil is not available. The related feathering pump advisory light, adjacent to the AUTOFEATHER switchlight, comes on during automatic operation of the auxiliary-feathering pump.

Once an autofeather is initiated, the system locks automatically to prevent both propellers autofeathering at the same time, and the ARM light goes out.

The autofeather system can be disarmed by:
• Pushing off the auto feather switchlight
• Retarding one or both POWER levers to FLT IDLE
• Both engine torque levers dropping below approximately 38%.

The Np underspeed cancel signal prevents the ECU from raising NH (if the engine is running in the case of an unscheduled feather command) in an attempt to maintain propeller rpm as the feathered propeller decreases below 780 rpm. If the underspeed-governing signal is not cancelled, the condition described above results in overtorquing of the feathering engine.

Autofeather test switches are used to test the system.

Propeller manual feathering is used during engine shutdown by moving the related condition lever to START & FEATHER. This system can also be used if the autofeather system malfunctions.

An alternate feather and unfeather system permits manual selection of either operation, if the condition lever selection fails to feather or unfeather the propeller. A three-position switch, on the autofeather control panel, controls the system.

Condition lever limit switches prevent selection of alternate feather or unfeather, unless the condition lever is in either the START & FEATHER or FUEL OFF position.

Pressurization of the airplane is dependent on three factors: (1) a positive, controlled inflow provided by the bleed-air system, (2) the flight compartment and cabin area being appropriately sealed, and (3) a controlled rate of air escaping from the fuselage.

The engine bleed air is supplied to and distributed by the air-conditioning system. Three outflow valves control the outflow air.

The outflow valves are controlled from the CABIN ALTITUDE control panel, on the overhead panel. There are independent controls and indicators to operate and monitor the system. Two aft outflow valves are located on the aft pressure dome. A safety outflow valve is located on the forward pressure bulkhead.

If cabin altitude is too high, a flight compartment warning light flashes.

The pressurization system can be controlled in one of four modes:
• Automatic
• Semi-automatic
• Manual
• Dump

With the system in AUTO mode, a preprogrammed cabin altitude computer does all pressure scheduling from takeoff to landing with minimal crew input. The computer receives inputs set by the crew and opens or closes the two aft outflow valves. This keeps a fixed schedule of cabin altitude versus airplane altitude for complete regulation.

When electrical power is first supplied, a self-test of the automatic system, illuminats the FAULT light for approximately two seconds. If no fauit is detected the light goes out.

When the aircraft is on the ground with the POWER levers at FLT IDLE or low power settings, the aft outflow valves are fully open to prevent pressurization.

At takeoff with bleed air on. When the POWER levers are advanced for takeoff. the cabin altitude descends to 140 feet below the existing cabin altitude until lift off and the landing gear is selected UP. At that time normal scheduling begins.

The flight sequence is started after takeoff or when bleeds are selected on, with AUTO mode selected. During this sequence, the cabin pressurization is computer controlled 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, the RATE knob can be adjusted to achieve a higher descent rate of cabin altitude.

If the set altitude on the cabin altitude indicator is less than the destination field elevation, the airplane will land pressurized. In this case, on landing, the cabin altitude will descend to field elevation at a pre-selected value on the rate selector for one minute, after which, the computer signals the aft outflow valves fully open. The maximum differential pressure for landing is 0.5 psi; this should be verified by the flight crew at the appropriate time.


Pressurization may be controlled semi-automatically if the automatic mode has failed or if direct control of cabin altitude is desired.

With the system in Semi-Automatic mode, part of the computer scheduling is bypassed and replaced by additional pilot input (e.g. Cabin cruising altitude, and rate of change of cabin altitude). Selecting the CAB SET/NORM switch to CAB SET will allow direct control of cabin altitude. The ALT and RATE knobs can be adjusted as necessary to give the desired altitude (Figure 18-10).

Pressurization may be controlled in MAN mode if the CPC fails or loses power, or the FAULT light comes on. The manual system consists of:
• cabin altitude MAN control knob
• Forward outflow valve

Electrical power is not required for operation in MAN mode.

Manual mode may be used with the automatic system operating during flight to evacuate smoke from the flight compartment, without depressurizing the airplane. Pressurization can be controlled through the forward outflow valve, when the AUTO/MAN/DUMP switch is set to AUTO.

Turning the cabin altitude manual control knob clockwise to INCR, allows suction caused by the slipstream to open the forward outflow valve. The control needle valve is mechanically connected to the outer dome of the front forward outflow valve and to ambient.

An arrow on the panel indicates that a clockwise selection opens the needle valve, venting the outer dome and increasing the cabin altitude. Full counterclockwise rotation closes the needle valve decreasing cabin altitude. The automatic system, trying to keep cabin pressure, will begin to close the aft outflow valve. This exhausts air and smoke through the forward outflow valve.

When operating in manual mode, the cabin altimeter, cabin differential pressure and cabin altitude rate of climb indicator should be monitored carefully.

The fast depressurization function may be done in the automatic and the manual modes. The AUTO/MAN/DUMP switch set to DUMP fully opens both aft outflow valves, preventing the aircraft from pressurizing. DUMP mode may be used for maximum smoke evacuation. The safety outflow valve is also used for ranl air ventilation during unpressurized flight.

Suction is generated by a venturi-ejector on the outside of the aft pressure dome adjacent to the valve. The ejector receives regulated 18 psi air from the de-icing system. A suction line runs from the ejector to the aft: outflow valves through a metering valve which regulates the degree of suction applied to the valves.

The metering valve is positioned electrically in response to the command signals from the cabin pressure controller to allow more or less suction to be applied. thus varying the pressure differential across the diaphragm and driving the aft outflow valves more open or closed.

The aft outflow valves are diaphragm-operated poppet valves. The valves are pneumatically opened by suction and spring-loaded closed. An electrical signal from the pressure controller creates a pressure differential within the valves.

If the cabin altitude is less than selected altitude, the valve opens to bleed some of the cabin pressure, increasing the cabin altitude. If the cabin altitude is more than the selected altitude, the outflow valves moves toward the closed position to restore cabin pressure (decrease cabin altitude).

During unpressurized flight the cabin and flight compartment can be ventilated with outside ram air. Without bleed air the ram air enters through the dorsal fin NACA vent, through a check valve and into the air conditioning ducting downstream of the pack. The ram air ventilates the cabin and flight compartment and then exhausts through the forward outflow valve.

The normal outflow valve is closed when the cabin altitude mode selector is in the MAN position.

The maximum differential pressure permitted by the controller is 5.5 ± 0.3 psi, which gives a cabin altitude of 8,000 feet at 25,000 feet ambient altitude. If differential pressure is more than 5.8 ± 0.15 psi, a pressure limiter in the forward and two aft valves opens to release the pressure. All safety valves also have a negative pressure relief valve, which will operate at - 0.1 psi differential to prevent external atmospheric pressure from being more than internal cabin pressure.

A FAULT light on the pressurization control panel will come on to show a system malfunction, or for two seconds during system self test.

A CABIN PRESS warning light will come on if cabin altitude is more than 10,000 feet.