Systems Flashcards
Explain the componentry of the fuel system
The airplane fuel system consists of two vented, integral fuel tanks with shutoff valves, a fuel selectors off warning system, a fuel reservoir, an ejector fuel pump, an electric auxiliary boost pump, a reservoir manifold assembly, a firewall shutoff valve, a fuel filter, an oil-to-fuel heater, an engine-driven fuel pump, a fuel control unit, a flow divider, dual manifolds, and 14 fuel nozzle assemblies. A fuel drain can and drain is also provided.
Explain the flow of the fuel in the fuel system
Fuel flows from the tanks through the two fuel tank shutoff valves at each tank. The fuel tank shutoff valves are mechanically controlled by two fuel selectors. Fuel flows by gravity from the shutoff valves in each tank to the fuel reservoir. The reservoir is located at the low point in the fuel system which maintains a head of fuel around the ejector boost pump and auxiliary boost pump which are contained within the reservoir. This head of fuel prevents pump cavitation in low-fuel quantity situations, especially during inflight manoeuvring. Fuel in the reservoir is pumped by the ejector boost pump or by the electric auxiliary boost pump to the reservoir manifold assembly. The ejector boost pump, which is driven by motive fuel flow from the FCU, normally provides fuel flow when the engine is operating.
In the event of failure of the ejector boost pump, the electric boost pump will automatically turn on, thereby supplying fuel flow to the engine. The auxiliary boost pump is also used to supply fuel flow during starting. Fuel in the reservoir manifold then flows through a fuel shutoff valve located on the aft side of the firewall. This shutoff valve enables the pilot to cut off all fuel to the engine. After passing through the shutoff valve, fuel is routed through a fuel filter located on the front side of the firewall. The fuel filter incorporates a bypass feature which allows fuel to bypass the filter in the event the filter becomes blocked with foreign material. A red filter bypass flag on the top of the filter extends upward when the filter is bypassing fuel. Fuel from the filter is then routed through the oil-to-fuel heater to the engine-driven fuel pump where fuel is delivered under pressure to the FCU.
The FCU meters the fuel and directs it to the flow divider which distributes the fuel to dual manifolds and 14 fuel nozzles located in the combustion chamber. Fuel rejected by the engine on shutdown drains into a fireproof fuel can located on the front left side of the firewall. The can should be drained during preflight inspection. If left unattended, the drain can fuel will overflow overboard. Fuel system venting is essential to system operation. Complete blockage of the vent system will result in decreased fuel flow and eventual engine stoppage. Venting is accomplished by check valve equipped vent lines, one from each fuel tank, which protrude from the trailing edge of the wing at the wing tips. Also the fuel reservoir is vented to both wing tanks.
Describe the airframe
The airplane is an all-metal, high-wing, single-engine airplane equipped with tricycle landing gear and designed for general utility purposes. The construction of the fuselage is a conventional formed sheet metal bulkhead, stringer, and skin design referred to as semimonocoque. Major items of structure are the front and rear carry-through spars to which the wings are attached, a bulkhead and forgings for main landing gear attachment and a bulkhead with attaching plates at its base for the strut-to-fuselage attachment of the wing struts. The externally braced wings, having integral fuel tanks, are constructed of a front and rear spar with formed sheet metal ribs, doublers, and stringers. The entire structure is covered with aluminium skin. The front spars are equipped with wing-to-fuselage and wing-to-strut attach fittings. The aft spars are equipped with wing-to-fuselage attach fittings.
The integral fuel tanks are formed by the front and rear spars, upper and lower skins, and inboard and outboard closeout ribs. Extensive use of bonding is employed in the fuel tank area to reduce fuelled tank sealing. Round-nosed ailerons and single-slot type flaps are of conventional formed sheet metal of each flap, is of conventional construction. The left aileron incorporates a servo tab while the right aileron incorporates a trimmable servo tab, both mounted on the outboard end of the aileron trailing edge.
The empennage consists of a conventional vertical stabiliser, rudder, horizontal stabiliser, and elevator. The vertical stabiliser consists of a forward and aft spar, sheet metal ribs and reinforcements, four skin panels, formed leading edge skins, and a dorsal fin. The rudder is constructed of a forward and aft spar, formed sheet metal ribs and reinforcements, and a wrap-around skin panel. The top of the rudder incorporates a leading edge extension which contains a balance weight. The horizontal stabiliser is constructed of a forward and aft spar, ribs and stiffeners, four upper and four lower skin panels, and two left and two right wrap-around skin panels which also form the leading edges. The horizontal stabiliser also contains dual jack screw type actuators for the elevator trim tabs. Construction of the elevator consists of a forward and aft spar, sheet metal ribs, upper and lower skin panels, and wrap-around skin panels for the leading and trailing edges.
An elevator trim tab is attached to the trailing edge of each elevator by full length piano-type hinges. Dual pushrods from each actuator located in the horizontal stabiliser transmit actuator movement to dual horns on each elevator trim tab to provide tab movement. Both elevator tip leading edge extensions provide aerodynamic balance and incorporate balance weights. A row of vortex generators on the top of the horizontal stabilise just forward of the elevator enhances nose down elevator and trim authority. To assure extended service life of the airplane, the entire airframe is corrosion proofed. Internally, all assemblies and sub-assemblies are coated with a chemical film conversion coating and are then epoxy primed. Steel parts in contact with aluminium structure are given a chromate dip before assembly. Externally, the complete airframe is painted with an overall coat of polyurethane paint which enhances resistance to corrosive elements in the atmosphere. Also, all control cables for the flight control system are of stainless steel construction.
Explain the flight controls
The airplane’s flight control system consists of conventional aileron, elevator and rudder control surfaces and a pair of spoilers mounted above the outboard ends of the flaps. The control surfaces are manually operated through mechanical linkage using a control wheel for the ailerons, spoilers and elevator and rudder/brake pedals for the rudder. The wing spoilers improve lateral control of the airplane at low speeds by disrupting lift over the appropriate flap. The spoilers are interconnected with the aileron system through a push-rod mounted to an arm on the aileron bellcrank. Spoiler travel is proportional to aileron travel for aileron deflections in excess of 5 degrees up. The spoilers are retracted throughout the remainder of aileron travel. Aileron servo tabs provide reduced manoeuvring control wheel forces. Fences on ailerons enhance lateral stability.
Explain the aircraft’s trim systems
Manually-operated aileron, elevator, and rudder trim systems are provided. Aileron trimming is achieved by a trimmable servo tab attached to the right aileron and connected mechanically to a knob located on the control pedestal. Elevator trimming is accomplished through two elevator trim tabs by utilising the vertically mounted trim control wheel on the top left side of the control pedestal. The airplane may also be equipped with an electric elevator trim system. Rudder trimming is accomplished through the nose wheel steering bungee connected to the rudder control system and a trim control wheel mounted on the control pedestal by rotating the horizontally mounted trim control wheel either left or right to the desired trim position.
Explain ground control of the aircraft
Effective ground control while taxiing is accomplished through nose wheel steering by using the rudder pedals. When a rudder pedal is depressed, a spring-loaded steering bungee (connected to the nose gear and to the rudder bars) will turn the nose wheel through an arc of approximately 15 degrees each side of centre. By applying either left or right brake, the degree of turn may be increased up to 56 degrees each side of centre.
Explain the wing flap system
The wing flaps are large span, single-slot type and are driven by an electric motor. Up to 30 degrees deflection. Mechanical stops at 10 and 20 degrees. A scale and white-tipped pointer on the left side of the selector lever provides a flap position indication. The system is protected by a pull-off type circuit breaker labelled FLAP MOTOR. Standby system can be used to operate the flaps in the event of primary system malfunction. Consists of standby motor, a guarded standby flap motor switch and a guarded standby flap motor up/down switch. The guarded standby flap motor switch has NORM and STBY positions. The NORM position permits operation of the flaps using the control pedestal mounted selector. STBY position is used to disable the dynamic braking of the primary flap motor when the standby flap motor system is operated. The standby flap motor up/down switch has UP, centre OFF and DOWN positions. To operate the flaps with the standby system, place the standby flap motor switch in STBY position. Then actuate the standby flap motor up/down switch momentarily to UP or DOWN and monitor the flap position indicator to obtain the desired flap position. Since the standby flap system does not have limit switches, the up/down switch should be terminated before the flaps reach full up or down travel. Standby flap system is protected by a pull-off type circuit breaker, labelled STBY FLAP MOTOR.
Explain the landing gear system
Tricycle type with a steerable nose wheel and two main wheels. Shock absorption is provided by the tubular spring-steel main landing gear struts, an interconnecting spring-steel tube between the two main landing gear struts, and the nose gear oil-filled shock strut and spring-steel drag link. Each main gear wheel is equipped with a hydraulically actuated single-disc brake on the inboard side of each wheel.
Explain the aircraft’s powerplant
The powerplant is a PT6A-114A free-turbine engine. It utilises two independent turbines; one driving a compressor in the gas generator section, and the second driving a reduction gearing for the propeller. Inlet air enters the engine through an annular plenum chamber formed by the compressor inlet case where it is directed to the compressor. The compressor consists of three axial stages combined with a single centrifugal stage, assembled as an integral unit. It provides a compression ratio of 7.0:1. A row of stator vanes located between each stage of compressor rotor blades diffuses the air, raises its static pressure and directs it to the next stage of compressor rotor blades. The compressed air passes through diffuser ducts which turn it 90 degrees in direction. It is then routed through straightening vanes into the combustion chamber. The combustion chamber liner located in the gas generator case consists of an annular reverse-flow weldment provided with varying sized perforations which allow entry of compressed air. The flow of air changes direction to enter the combustion chamber liner where it reverses direction and mixes with fuel. The location of the combustion chamber liner eliminates the need for a long shaft between the compressor and the compressor turbine, thus reducing the overall length and weight of the engine. Fuel is injected into the combustion chamber liner by 14 simplex nozzles supplied by a dual manifold. The mixture is initially ignited by two spark igniters which protrude into the combustion chamber liner. The resultant gases expand from the combustion chamber liner, reverse direction and pass through the compressor turbine guide vanes to the compressor turbine. The turbine guide vanes ensure that the expanding gases impinge on the turbine blades at the proper angle, with a minimum loss of energy. The still expanding gases pass forward through a second set of stationary guide vanes to drive the power turbine. The compressor and power turbines are located in the approximate centre of the engine with their shafts extending in opposite directions. The exhaust gas from the power turbine is directed through an exhaust plenum to the atmosphere via a single exhaust port on the right side of the engine. The engine is flat rated at 675 shaft horsepower (1865 foot-pounds torque at 1900 RPM varying linearly to 1970 foot-pounds torque at 1800 RPM; below 1800 RPM, the maximum torque value remains constant at 1970 foot-pounds). Between 1800 and 1600 prop RPM, the gearbox torque limit of 1970 foot-pounds will not allow the full flat rating of 675 SHP to be achieved. The speed of the gas generator (compressor) turbine (Ng) is 37500 RPM at 100 Ng. Maximum permissible speed of the gas generator is 38100 RPM which equals 101.6 Ng. The power turbine speed is 33000 RPM at a propeller shaft speed of 1900 RPM (a reduction ratio of 0.0576:1). All engine-driven accessories, with the exception of the propeller tachometer-generator and the propeller governors, are mounted on the accessory gearbox located at the rear of the engine. These are driven by the compressor turbine with a coupling shaft which extends the drive through a conical tube in the oil tank centre section. The engine oil supply is contained in an integral tank which forms part of the compressor inlet case. The tank has a capacity of 9.5 US quarts and is provided with a dipstick and drain plug. The power turbine drives the propeller through a two-stage planetary reduction gearbox located on the front of the engine. The gearbox embodies an integral torquemeter device which is instrumented to provide an accurate indication of the engine power output.
Explain the Power Lever
The power lever is connected through linkage to a cam assembly mounted in front of the fuel control unit at the rear of the engine. The power lever controls engine power through the full range from maximum takeoff power back through idle to full reverse. The lever also selects propeller pitch when in the BETA range. The power lever has MAX, IDLE, and BETA and REVERSE range positions. The range from MAX position through IDLE enables the pilot to select the desired power output from the engine. The BETA range enables the pilot to control propeller blade pitch from idle thrust back through a zero or no-thrust condition to maximum reverse thrust. Caution: The propeller reversing linkage can be damaged if the power lever is moved aft of the IDLE position when the propeller is feathered.
Explain the Emergency Power Lever
The emergency power lever is connected through linkage to the manual override lever on the fuel control unit and governs fuel supply to the engine should a pneumatic malfunction occur in the fuel control unit. When the engine is operating, a failure of any pneumatic signal input to the fuel control unit will result in the fuel flow decreasing to minimum idle (about 48% Ng at sea level and increasing with altitude). The emergency power lever allows the pilot to restore power in the event of such a failure. The emergency power lever has NORMAL, IDLE, and MAX positions. The NORMAL position is used for all normal engine operation when the fuel control unit is operating normally and engine power is selected by the power lever. The range from IDLE position to MAX governs engine power and is used when a pneumatic malfunction has occurred in the fuel control unit and the power lever is ineffective. A mechanical stop in the lever slot requires that the emergency power lever be moved to the left to clear the stop before it can be moved from the NORMAL (full aft) position to the IDLE position. Note: The knob on the emergency power lever has crosshatching which is visible when the lever is in MAX position. Also, the emergency power lever is annunciated on the annunciator panel whenever it is unstowed from the NORMAL position. These precautions are intended to preclude starting of the engine with the emergency power lever inadvertently placed in any position other than NORMAL.
Explain the Propeller Control Lever
The propeller control lever is connected through linkage to the propeller governor mounted on top of the front section of the engine and controls propeller governor settings from the maximum RPM position to full feather. The propeller control lever has MAX, MIN, and FEATHER positions. The MAX position is used when high RPM is desired and governs the propeller speed at 1900 RPM. Propeller control lever settings from the MAX position to MIN permit the pilot to select the desired engine RPM for cruise. The FEATHER position is used during normal engine shutdown to stop rotation of the power turbine and front section of the engine. Since lubrication is not available after the gas generator section if the engine has shut down, rotation of the forward section of the engine is not desirable. Also, feathering the propeller when the engine is shut down minimises propeller windmilling during windy conditions. A mechanical stop in the lever slot requires that the propeller control lever be moved to the left to clear the stop before it can be moved into or out of the FEATHER position.
Explain the Fuel Condition Lever
The fuel condition lever is connected through linkage to a combined lever and stop mechanism on the fuel control unit. The lever and stop also function as an idle stop for the fuel control unit rod. The fuel condition lever controls the minimum RPM of the gas generator turbine (Ng) when the power lever is in the IDLE position. The fuel condition lever has CUTOFF, LOW IDLE, and HIGH IDLE positions. The CUTOFF position shuts off all fuel to the engine fuel nozzles. LOW IDLE positions the control rod stop to provide an RPM of 52% Ng. HIGH IDLE positions the control rod stop to provide an RPM of 65% Ng.
Explain the Torque Indicator
The torque indicator is located on the upper portion of the instrument panel and indicates the torque being produced by the engine. The transmitter senses the difference between the engine torque pressure and the pressure in the engine case and transmits this data to the torque indicator. The torque indicator converts this information into an indication of torque in foot-pounds. The torque indicator system is powered by 28-volt DC power through a circuit breaker, labelled TRQ IND, on the left sidewall switch and circuit breaker panel. On other Cargo Versions and the Passenger Version, the torque indicator is pressure actuated. Two independent lines enter the back of the torque indicator. One line measures the engine torque pressure and one line measures the reduction gearbox internal pressure. The torque indicator monitors the engine torque pressure and converts this pressure into an indication of torque in foot-pounds. Instrument markings indicate that the normal operating range (green arc) is from 0 to 1865 foot-pounds, the alternate power range (striped green arc) is from 1865 to 1970 foot-pounds, and maximum torque (red line) is 1970 foot-pounds. Maximum takeoff torque is denoted by ‘‘1.0.’’ and a red wedge at 1865 foot-pounds.
Explain the Propeller RPM Indicator
The propeller RPM indicator is located on the upper portion of the instrument panel. The instrument is marked in increments of 50 RPM and indicates propeller speed in revolutions per minute. The instrument is electrically-operated from the propeller tachometer-generator which is mounted on the right side of the front case. Instrument markings indicate a normal operating range (green arc) of from 1600 to 1900 RPM and a maximum (red line) of 1900 RPM.
Explain the ITT Indicator
The ITT (interturbine temperature) indicator is located on the upper portion of the instrument panel. The instrument displays the gas temperature between the compressor and power turbines. Instrument markings indicate a normal operating range (green arc) of from 100°C to 740°C, and a maximum (red line) of 805°C. Also, instrument markings indicate a maximum starting temperature (red triangle) of 1090°C.
Explain the Ng% RPM Indicator
The Ng% RPM indicator is located on the upper portion of the instrument panel. The instrument indicates the percent of gas generator RPM based on a figure of 100% at 37,500 RPM. The instrument is electrically-operated from the gas generator tachometer-generator mounted on the lower right-hand portion of the accessory case. Instrument markings indicate a normal operating range (green arc) of from 52% to 101 .6% and a maximum (red line) of 101.6%.
Explain the Oil Pressure Gauge
The oil pressure gage is the left half of a dual-indicating instrument I mounted on the upper portion of the instrument panel. A direct pressure oil line from the engine delivers oil at engine operating pressure to the oil pressure gage.
Explain the Oil Temperature Gauge
The oil temperature gage is the right half of a dual-indicating I instrument mounted on the upper portion of the instrument panel. The instrument is operated by an electrical-resistance type temperature sensor which receives power from the airplane electrical system.
Explain the Engine Lubrication System
The lubrication system consists of a pressure system, a scavenge system and a breather system. The main components of the lubrication system include an integral oil tank at the back of the engine, an oil pressure pump at the bottom of the oil tank, an external double-element scavenge pump located on the back of the accessory case, an internal double-element scavenge pump located inside the accessory gearbox, an oil-to-fuel heater located on the top rear of the accessory case, an oil filter located internally on the right side of the oil tank and an oil cooler located on the right side of the nose cowl. A large capacity oil cooler is installed to increase the hot day outside air temperature limits for flight operations. Oil is drawn from the bottom of the oil tank through a filter screen where it passes through a pressure relief valve for regulation of oil pressure. The pressure oil is then delivered from the main oil pump to the oil filter where extraneous matter is removed from the oil and precluded from further circulation.Pressure oil is then routed through passageways to the engine bearings, reduction gears, accessory drives, torque meter and propeller governor. Also, pressure oil is routed to the oil-to-fuel heater where it then returns to the oil tank. After cooling and lubricating the engine moving parts, oil is scavenged as follows: Oil from the number 1 bearing compartment is returned by gravity into the accessory gearbox. Oil from the number 2 bearing is scavenged by the front element of the internal scavenge pump back into the accessory gearbox. Oil from the number 3 and number 4 bearings is scavenged by the front element of the external scavenge pump into the accessory gearbox. Oil from the propeller governor, front thrust bearing, reduction gear accessory drives, and torquemeter is scavenged by the rear element of the external scavenge pump where it is routed through a thermostatically-controlled oil cooler and then returned to the oil tank. Also, the rear element of the internal scavenge pump scavenges oil from the accessory case and routes it through the oil cooler where it then returns to the oil tank.
Breather air from the engine bearing compartments and from the accessory and reduction gearboxes is vented overboard through a centrifugal breather installed in the accessory gearbox. The bearing compartments are connected to the accessory gearbox by cored passages and existing scavenge oil return lines. A bypass valve immediately upstream of the front element of the internal scavenge pump vents the accessory gearbox when the engine is operating at high power. An oil dipstick/filler cap is located at the rear of the engine on the left side and is accessible when the left side of the upper cowling is raised. Markings which indicate U.S. quarts low if the oil is hot are provided on the dipstick to facilitate oil servicing. The oil tank capacity is 9.5 U.S. quarts and total system capacity is 14 U.S. quarts.
Explain the Ignition System
The ignition system consists of two igniters, an ignition exciter, two high-tension leads, an ignition monitor light, an ignition switch, and a starter switch. Engine ignition is provided by two igniters in the engine combustion chamber. The igniters are energised by the ignition exciter mounted on the engine mount on the right side of the engine compartment. Electrical energy from the ignition exciter is transmitted through two high-tension leads to the igniters in the engine. The ignition system is normally energised only during engine start.
Ignition is controlled by an ignition switch and a starter switch located on the left sidewall switch and circuit breaker panel. The ignition switch has two positions, ON and NORMAL. The NORMAL position of the switch arms the ignition system so that ignition will be obtained when the starter switch is placed in the START position. The NORMAL position is used during all ground starts and during air starts with starter assist. The ON position of the switch provides continuous ignition regardless of the position of the starter switch. This position is used for air starts without starter assist, for operation on water or slush-covered runways, during flight in heavy precipitation, during inadvertent icing encounters until the inertial separator has been in bypass for 5 minutes, and when near fuel exhaustion as indicated by illumination of the RESERVOIR FUEL LOW annunciator.
The main function of the starter switch is control of the starter for rotating the gas generator portion of the engine during starting. However, it also provides ignition during starting. For purposes of this discussion, only the ignition functions of the switch are described. The starter switch has three positions, OFF, START, and MOTOR. The OFF position shuts off the ignition system and is the normal position at all times except during engine start or engine clearing. The START position energises the engine ignition system provided the ignition switch is in the NORMAL position. After the engine has started during a ground or air start, the starter switch must be manually positioned to OFF for generator operation.
A green annunciator, located on the annunciator panel, is labelled IGNITION ON, and will illuminate when electrical power is being applied to the igniters. The ignition system is protected by a pull-off type circuit breaker, labelled IGN.
Explain the Air Induction System
The engine air inlet is located at the front of the engine nacelle to the left of the propeller spinner. Ram air entering the inlet flows through ducting and an inertial separator system and then enters the engine through a circular plenum chamber where it is directed to the compressor by guide vanes. The compressor air inlet incorporates a screen which will prevent entry of large articles, but does not filter the inlet air.
Explain the Inertial Separator System
An inertial separator system in the engine air inlet duct prevents moisture particles from entering the compressor air inlet plenum when in bypass mode. The inertial separator consists of two movable vanes and a fixed airfoil which, during normal operation, route the inlet air through a gentle turn into the compressor air inlet plenum. When separation of moisture particles is desired, the vanes are positioned so that the inlet air is forced to execute a sharp turn in order to enter the inlet plenum. This sharp turn causes any moisture particles to separate from the inlet air and discharge overboard through the inertial separator outlet in the left side of the cowling.
Inertial separator operation is controlled by a T-handle located on the lower instrument panel. The T-handle is labelled BYPASS-PULL, NORMAL-PUSH. The inertial separator control should be moved to the BYPASS position prior to running the engine during ground or flight operation in visible moisture (clouds, rain, snow, ice crystals) with an OAT of 4°C or less. It may also be used for ground operations or take-off’s from dusty, sandy field conditions to minimise ingestion of foreign particles into the compressor. The normal position is used for all other operations. The T-handle locks in the NORMAL position by rotating the handle clockwise 1/4 turn to its vertical position. To unlock, push forward slightly and rotate the handle 90° counter-clockwise. The handle can then be pulled into the BYPASS position. Once moved to the BYPASS position, air loads on the movable vanes hold them in this position.
NOTE: When moving the inertial separator control from bypass to normal position during flight, reduction of engine power will reduce the control forces.
Explain the Exhaust System
The exhaust system consists of a primary exhaust pipe attached to the right side of the engine just aft of the propeller reduction gearbox. A secondary exhaust duct, fitted over the end of the primary exhaust pipe, carries the exhaust gases away from the cowling and into the slipstream. The juncture of the primary exhaust pipe and secondary exhaust duct is located directly behind the oil cooler. Since the secondary exhaust duct is of larger diameter than the primary exhaust pipe, a venturi effect is produced by the flow of exhaust. This venturi effect creates a suction behind the oil cooler which augments the flow of cooling air through the cooler. This additional airflow improves oil cooling during ground operation of the engine.
Explain the Engine Fuel System
The engine fuel system consists of an oil-to-fuel heater, an engine-driven fuel pump, a fuel control unit, a flow divider and dump valve, a dual fuel manifold with 14 simplex nozzles, and two fuel drain lines. The system provides fuel flow to satisfy the speed and power demands of the engine. Fuel from the airplane reservoir is delivered to the oil-to-fuel heater which is essentially a heat exchanger which utilises heat from the engine lubricating oil system to preheat the fuel in the fuel system. A fuel temperature-sensing oil bypass valve regulates the fuel temperature by either allowing oil to flow through the heater circuit or bypass it to the engine oil tank.
Fuel from the oil-to-fuel heater then enters the engine-driven fuel pump chamber through a 74-micron inlet screen. The inlet screen is spring-loaded and should it become blocked, the increase in differential pressure will overcome the spring and allow unfiltered fuel to flow into the pump chamber. The pump increases the fuel pressure and delivers it to the fuel control unit via a 10-micron filter in the pump outlet. A bypass valve and cored passages in the pump body enables unfiltered high pressure fuel to flow to the fuel control unit in the event the outlet filter becomes blocked.
The fuel control unit consists of a fuel metering section, a temperature compensating section, and a gas generator (Ng) pneumatic governor. The fuel control unit determines the proper fuel schedule to provide the power required as established by the power lever input. This is accomplished by controlling the speed of the compressor turbine. The temperature compensating section alters the acceleration fuel schedule to compensate for fuel density differences at different fuel temperatures, especially during engine start. The power turbine governor, located in the propeller governor housing, provides power turbine overspeed protection in the event of propeller governor failure. This is accomplished by limiting fuel to the gas generator. During reverse thrust operation, maximum power turbine speed is controlled by the power turbine governor. The temperature compensator alters the acceleration fuel schedule of the fuel control unit to compensate for variations in compressor inlet air temperature. Engine characteristics vary with changes in inlet air temperature, and the acceleration fuel schedule must, in turn, be altered to prevent compressor stall and/or excessive turbine temperatures.
The flow divider schedules the metered fuel, from the fuel control unit, between the primary and secondary fuel manifolds. The fuel manifold and nozzle assemblies deliver fuel to the combustion chamber through 10 primary and 4 secondary fuel nozzles. During engine start, metered fuel is delivered initially by the primary nozzles, with the secondary nozzles cutting in above a preset value. All nozzles are operative at idle and above. When the fuel cutoff valve in the fuel control unit closes during engine shutdown, both primary and secondary manifolds are connected to a dump valve port and residual fuel in the manifolds is allowed to drain into the fuel drain can attached to the firewall where it can be drained daily.
Explain the Cooling System
No external cooling provisions are provided for the PT6A-114A engine in this installation. However, the engine incorporates an extensive internal air system which provides for bearing compartment sealing and for compressor and power turbine disk cooling.
Explain the Starting System
The starting system consists of a starter/generator, a starter switch, and a starter annunciator light. The starter/generator functions as a motor for engine starting and will motor the gas generator section until a speed of 46% Ng is reached, at which time, the start cycle will automatically be terminated by a speed sensing switch located in the starter/generator. The starter/generator is controlled by a three-position starter switch located on the left sidewall switch and circuit breaker panel. The switch has OFF, START, and MOTOR positions. The OFF position de-energises the ignition and starter circuits and is the normal position at all times except during engine start. The START position of the switch energises the starter/generator which rotates the gas generator portion of the engine for starting. Also, the START position energises the ignition system, provided the ignition switch is in the NORMAL position. When the engine has started, the starter switch must be manually placed in the OFF position to de-energise the ignition system and activate the generator system. The MOTOR position of the switch motors the engine without having the ignition circuit energised and is used for motoring the engine when a start is not desired. This can be used for clearing fuel and engine start is not desired. This can be used for clearing fuel from the engine, washing the engine compressor, etc. The MOTOR position is spring-loaded back to the OFF position. Also, an interlock between the MOTOR position of the starter switch and the ignition switch prevents the starter from motoring unless the ignition switch is in the NORMAL position. This prevents unintentional motoring of the engine with the ignition on. Starter contactor operation is indicated by an amber annunciator, labelled STARTER ENERGISED, on the annunciator panel.
Explain the Oil Pump
Pressure oil is circulated from the integral oil tank through the engine lubrication system by a self-contained, gear-type pressure pump located in the lowest part of the oil tank. The oil pump is contained in a cast housing which is bolted to the front face of the accessory diaphragm, and is driven by the accessory gear shaft. The oil pump body incorporates a circular mounting boss to accommodate a check valve, located in the end of the filter housing. A second mounting boss on the pump accommodates a pressure relief valve.
Explain the Fuel Pump
The engine-driven fuel pump is mounted on the accessory gearbox at the 2 o’clock position. The pump is driven through a gear shaft and splined coupling. The coupling splines are lubricated by oil mist from the auxiliary gearbox through a hole in the gear shaft. Another splined coupling shaft extends the drive to the fuel control unit which is bolted to the rear face of the pump. Fuel from the oil-to-fuel heater enters the fuel pump through a 74-micron inlet screen. Then, fuel enters the pump gear chamber, is boosted to high pressure, and delivered to the fuel control unit through a 10-micron pump outlet filter. A bypass valve and cored passages in the pump casing enable unfiltered high pressure fuel to flow from the pump gears to the fuel control unit should the outlet filter become blocked. An internal passage originating at the mating face with the fuel control unit returns bypass fuel from the fuel control unit to the pump inlet downstream of the inlet screen. A pressure regulating valve in this line serves to pressurise the pump gear bushings.
Explain the Ng Tachometer-Generator
The Ng tachometer-generator produces an electric current which is used In conjunction with the gas generator RPM Indicator to indicate gas generator RPM. The Ng tachometer-generator drive and mount pad is located at the 5 o’clock position on the accessory gearbox and is driven from the internal scavenge pump. Rotation is counter-clockwise with a drive ratio of 0.1121 :1.
Explain the Propeller Tachometer-Generator
The propeller tachometer-generator produces an electric current which is used in conjunction with the propeller RPM indicator. The propeller tachometer-generator drive and mount pad is located on the right side of the reduction gearbox case and rotates clockwise with a drive ratio of 0.1273:1.
Explain the Torquemeter
The torquemeter is a hydro-mechanical torque measuring device located inside the first stage reduction gear housing to provide an accurate indication of engine power output. The difference between the torquemeter pressure and the reduction gearbox internal pressure accurately indicates the torque being produced. The two pressures are internally routed to bosses located on the top of the reduction gearbox front case and are then plumbed to the torquemeter indicator which indicates the correct torque pressure.
Explain the Starter/Generator
The starter/generator is mounted on the top of the accessory case at the rear of the engine. The starter/generator is a 28-volt, 200-amp engine-driven unit that functions as a motor for engine starting and, after engine start, as a generator for the airplane electrical system. When operating as a starter, a speed sensing switch in the starter/generator will automatically shut down the starter, thereby providing overspeed protection and automatic shutoff. The starter/generator is air cooled by an integral fan and by ram air ducted from the front of the engine cowling.
Explain the Interturbine Temperature Sensing System
The interturbine temperature sensing system is designed to provide the operator with an accurate indication of engine operating temperatures taken between the compressor and power turbines. The system consists of twin leads, two bus bars, and eight individual chromel-alumel thermocouple probes connected in parallel. Each probe protrudes through a threaded boss on the power turbine stator housing into an area adjacent to the leading edge of the power turbine vanes. The probe is secured to the boss by means of a floating, threaded fitting which is part of the thermocouple probe assembly. Shielded leads connect each bus bar assembly to a terminal block which provides a connecting point for external leads to the ITT indicator in the airplane cabin.
Explain the Propeller Governor
The propeller governor is located in the 12 o’clock position on the front case of the reduction gearbox. Under normal conditions, the governor acts as a constant speed unit, maintaining the propeller speed selected by the pilot by varying the propeller blade pitch to match the load to the engine torque. The propeller governor also has a power turbine governor section built into the unit. Its function is to protect the engine against a possible power turbine overspeed in the event of a propeller governor failure.
If such an overspeed should occur, a governing orifice in the propeller governor is opened by flyweight action to bleed off compressor discharge pressure through the governor and computing section of the fuel control unit. When this occurs, compressor discharge pressure, acting on the fuel control unit governor bellows, decreases and moves the metering valve in a closing direction, thus reducing fuel flow to the flow divider.
