TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN

TRAINING HANDOUT PT REVEN GLOBAL AIRTRANSPOR

OPERATION DEPT. SEPTEMBER 2021

TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN TABLE OF CONTENT

Contents AIRFRAME ................................................................................................................................................................... 1 GENERAL .................................................................................................................................................................. 1 FUSELAGE ................................................................................................................................................................. 1 WINGS ........................................................................................................................................................................ 1 EMPENNAGE............................................................................................................................................................. 2 CARGO POD .............................................................................................................................................................. 2 FLIGHT CONTROLS .................................................................................................................................................. 1 AILERONS ................................................................................................................................................................. 1 SPOILERS ................................................................................................................................................................... 1 ELEVATORS .............................................................................................................................................................. 1 RUDDER ..................................................................................................................................................................... 1 CONTROL LOCKS .................................................................................................................................................... 2 WING FLAP SYSTEM ............................................................................................................................................... 2 STANDBY FLAP SYSTEM ....................................................................................................................................... 2 COCKPIT ARRANGEMENT ..................................................................................................................................... 1 INSTRUMENT PANEL .............................................................................................................................................. 1 CONTROL PEDESTAL ............................................................................................................................................. 1 LEFT SIDEWALL SWITCH AND CIRCUIT BREAKER PANEL .......................................................................... 1 OVERHEAD PANEL ................................................................................................................................................. 1 PILOTS SEATS .......................................................................................................................................................... 2 CABIN DOORS ............................................................................................................................................................. 1 CABIN ENTRY DOORS ............................................................................................................................................ 1 CREW ENTRY DOORS ............................................................................................................................................. 1 PASSENGER ENTRY DOOR .................................................................................................................................... 1 CARGO DOORS ......................................................................................................................................................... 1 WARNING SYSTEMS ................................................................................................................................................. 1 ANNUNCIATOR PANEL .......................................................................................................................................... 1 FUEL SELECTORS OF WARNING SYSTEM ......................................................................................................... 4 STALL WARNING SYSTEM .................................................................................................................................... 4 OVERSPEED WARNING SYSTEM ......................................................................................................................... 4 LANDING GEAR AND BRAKE SYSTEM ................................................................................................................ 1 LANDING GEAR ....................................................................................................................................................... 1 BRAKE SYSTEM ....................................................................................................................................................... 1 TURBINE ENGINE THEORY .................................................................................................................................... 1 AIR INLET DUCT ...................................................................................................................................................... 1 COMPRESSORS......................................................................................................................................................... 1 THE CENTRIFUGAL COMPRESSOR ................................................................................................................. 1 THE AXIAL FLOW COMPRESSOR .................................................................................................................... 1 COMPRESSOR STALLS ....................................................................................................................................... 2 THE COMBUSTION CHAMBER ............................................................................................................................. 3

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN TABLE OF CONTENT TURBINE .................................................................................................................................................................... 3 THE EXHAUST .......................................................................................................................................................... 3 TURBOPROP ENGINE POWER CONVERSION .................................................................................................... 4 NEW

THE FIXED SHAFT DESIGN .................................................................................................................. 4

NEW

THE FREE TURBINE............................................................................................................................... 4

PT6 OPERATING PRINCIPLES ................................................................................................................................ 4 POWER PLANT............................................................................................................................................................ 1 NEW DESCRIPTION ................................................................................................................................................. 1 ENGINE CONTROLS ................................................................................................................................................ 1 POWER LEVER ..................................................................................................................................................... 1 EMERGENCY POWER LEVER ............................................................................................................................ 1 PROPELLER CONTROL LEVER ......................................................................................................................... 2 FUEL CONDITION LEVER .................................................................................................................................. 2 QUADRANT FRICTION LOCK ............................................................................................................................ 2 ENGINE INSTRUMENTS ......................................................................................................................................... 3 TORQUE INDICATOR .......................................................................................................................................... 3 PROPELLER RPM INDICATOR........................................................................................................................... 3 ITT INDICATOR .................................................................................................................................................... 3 Ng% RPM INDICATOR ......................................................................................................................................... 3 FUEL FLOW INDICATOR .................................................................................................................................... 3 OIL PRESSURE GAGE .......................................................................................................................................... 3 ENGINE LUBRICATION SYSTEM .......................................................................................................................... 5 IGNITION SYSTEM .................................................................................................................................................. 6 NORMAL POSITION ............................................................................................................................................. 6 ON POSITION ........................................................................................................................................................ 6 STARTING SYSTEM ................................................................................................................................................. 7 OFF POSITION ....................................................................................................................................................... 7 START POSITION ................................................................................................................................................. 7 MOTOR POSITION ................................................................................................................................................ 7 ENGINE START CONSIDERATIONS ..................................................................................................................... 8 AIR INDUCTION SYSTEM ...................................................................................................................................... 9 INERTIAL SEPARATOR SYSTEM ...................................................................................................................... 9 EXHAUST SYSTEM ................................................................................................................................................ 10 ENGINE FUEL SYSTEM ......................................................................................................................................... 11 OIL-FUEL HEATER............................................................................................................................................. 11 FUEL PUMP ......................................................................................................................................................... 11 FUEL CONTROL UNIT ....................................................................................................................................... 11 ENGINE ACCESSORIES ......................................................................................................................................... 16 OIL PUMP............................................................................................................................................................. 16 FUEL PUMP ......................................................................................................................................................... 16 Ng TACHO GENERATOR .................................................................................................................................. 16 TORQUE METER ................................................................................................................................................ 16

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN TABLE OF CONTENT STARTER/GENERATOR .................................................................................................................................... 16 ITT SENSING SYSTEM ...................................................................................................................................... 16 PROPELLER GOVERNOR .................................................................................................................................. 16 PROPELLER OVERSPEED GOVERNOR .......................................................................................................... 17 ENGINE FIRE DETECTION SYSTEM ............................................................................................................... 17 ENGINE REDUCTION GEAR SYSTEM ............................................................................................................ 17 CHIP DETECTORS .............................................................................................................................................. 17 OIL BREATHER DRAIN CAN............................................................................................................................ 17 PROPELLER ................................................................................................................................................................. 1 PROPELLER GOVERNOR ........................................................................................................................................ 1 CONSTANT SPEED MODE .................................................................................................................................. 1 OPERATION........................................................................................................................................................... 1 LOW PITCH STOP ................................................................................................................................................. 2 BETA MODE .......................................................................................................................................................... 2 REVERSE CONTROL ............................................................................................................................................ 2 FEATHERING ........................................................................................................................................................ 3 UNFEATHERING .................................................................................................................................................. 3 POWER TURBINE (Nf) OVERSPEED GOVERNOR .............................................................................................. 4 PROPELLER OVERSPEED GOVERNOR ............................................................................................................ 5 FUEL SYSTEM ............................................................................................................................................................. 1 COMPONENTS .......................................................................................................................................................... 1 SYSTEM ..................................................................................................................................................................... 1 FIREWALL SHUTOFF VALVE ................................................................................................................................ 2 FUEL TANK SELECTORS ........................................................................................................................................ 2 AUXILIARY BOOST PUMP SWITCH ..................................................................................................................... 2 OFF POSITION ....................................................................................................................................................... 2 NORM POSITION .................................................................................................................................................. 2 ON POSITION ........................................................................................................................................................ 2 FUEL FLOW INDICATOR ........................................................................................................................................ 3 FUEL QUANTITY INDICATORS............................................................................................................................. 4 ANNUNCIATORS ...................................................................................................................................................... 4 FUEL LOW ............................................................................................................................................................. 4 RESERVOIR FUEL LOW ...................................................................................................................................... 4 FUEL PRESSURE LOW ........................................................................................................................................ 4 AUX FUEL PUMP ON ........................................................................................................................................... 4 FUEL DRAIN VALVES ............................................................................................................................................. 4 FUEL DRAIN CAN .................................................................................................................................................... 4 FUEL PUMP DRAIN RESERVOIR ........................................................................................................................... 5 ELECTRICAL SYSTEM ............................................................................................................................................. 1 ELECTRICAL BUSES ............................................................................................................................................... 1 STARTER-GENERATOR .......................................................................................................................................... 1 GENERATOR CONTROL UNIT (GCU) ................................................................................................................... 1

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN TABLE OF CONTENT GROUND POWER MONITOR .................................................................................................................................. 1 GROUND SERVICE PLUG RECEPTACLE ............................................................................................................. 2 EXTERNAL POWER SWITCH ................................................................................................................................. 2 BATTERY ................................................................................................................................................................... 3 THERMAL RUNAWAY ........................................................................................................................................ 3 BATTERY HOT light ............................................................................................................................................. 3 BATTERY OVERHEAT light ................................................................................................................................ 3 BATTERY SWITCH .................................................................................................................................................. 4 GENERATOR SWITCH ............................................................................................................................................. 4 RESET AND TRIP .................................................................................................................................................. 4 ON POSITION ........................................................................................................................................................ 4 AVIONICS POWER SWITCHES .............................................................................................................................. 5 AVIONICS BUS TIE SWITCH .................................................................................................................................. 5 CIRCUIT BREAKERS ............................................................................................................................................... 5 VOLT/AMMETER ..................................................................................................................................................... 5 GEN POSITION ...................................................................................................................................................... 5 ALT POSITION ...................................................................................................................................................... 5 BATT POSITION .................................................................................................................................................... 5 VOLT POSITION ................................................................................................................................................... 5 STANDBY ELECTRICAL SYSTEM ........................................................................................................................ 6 LIGHTING .................................................................................................................................................................... 1 EXTERIOR LIGHTING ............................................................................................................................................. 1 NAVIGATION LIGHTS ......................................................................................................................................... 1 LANDING LIGHTS ................................................................................................................................................ 1 TAXI/RECOGNITION LIGHTS ............................................................................................................................ 1 STROBE LIGHTS ................................................................................................................................................... 1 FLASHING BEACON LIGHT ............................................................................................................................... 1 COURTESY LIGHTS ............................................................................................................................................. 1 INTERIOR LIGHTING............................................................................................................................................... 2 CONTROL KNOBS ................................................................................................................................................ 2 MAP LIGHT............................................................................................................................................................ 2 CABIN LIGHTS...................................................................................................................................................... 2 PASSENGER READING LIGHTS ........................................................................................................................ 2 NO SMOKING/FASTEN BELT SIGN .................................................................................................................. 2 AIR CONDITIONING .................................................................................................................................................. 1 VENTILATING AIR................................................................................................................................................... 1 VENTILATION FAN SYSTEM ............................................................................................................................. 1 INSTRUMENT PANEL VENT KNOBS ................................................................................................................ 1 CABIN HEATING SYSTEM ..................................................................................................................................... 2 BLEED AIR HEAT SWITCH ................................................................................................................................. 2 TEMPERATURE SELECTOR KNOB ................................................................................................................... 2 MIXING AIR PUSH-PULL CONTROL ................................................................................................................ 2

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN TABLE OF CONTENT AFT /FORWARD CABIN PUSH-PULL CONTROL ................................................................................................ 4 DEFROST/FORWARD CABIN PUSH-PULL CONTROL ....................................................................................... 4 CABIN HEAT FIREWALL SHUTOFF KNOB ......................................................................................................... 4 OXYGEN........................................................................................................................................................................ 1 INSTRUMENTS ............................................................................................................................................................ 1 PITOT-STATIC SYSTEM .......................................................................................................................................... 1 VACUUM SYSTEM ................................................................................................................................................... 1 SUCTION GAGE .................................................................................................................................................... 1 VACUUM-LOW WARNING ANNUNCIATOR ................................................................................................... 1 ELECTRICALLY-OPERATED GYROS ................................................................................................................... 2 ICE AND RAIN PROTECTION.................................................................................................................................. 1 PNEUMATIC DE-ICE BOOTS .................................................................................................................................. 1 PROPELLER ANTI-ICE BOOTS............................................................................................................................... 2 WINDSHIELD ANTI-ICE PANEL ............................................................................................................................ 2 ICE DETECTOR LIGHT ............................................................................................................................................ 2 AUTOFLIGHT SYSTEM AND AVIONICS .............................................................................................................. 1 THE KFC-150 AUTOPILOT SYSTEM ..................................................................................................................... 1 DEFINITIONS AND TERMS ................................................................................................................................ 1 THE AUTOPILOT AND FLIGHT DIRECTOR COMPUTER .............................................................................. 1 BEFORE TAKEOFF RELIABILITY TESTS ......................................................................................................... 4 AUTOPILOT LIMITATIONS ................................................................................................................................ 4 MAXIMUM ALTITUDE LOSS DUE TO MALFUNCTION ................................................................................ 4 ATTITUDE INDICATOR ........................................................................................................................................... 5 COMMAND BAR ................................................................................................................................................... 5 FCI SYMBOLIC AIRCRAFT ................................................................................................................................. 5 HORISONTAL SITUATION INDICATOR ............................................................................................................... 6 NAV FLAG ............................................................................................................................................................. 6 LUBBER LINE ....................................................................................................................................................... 6 HEADING WARNING FLAG (HDG) ................................................................................................................... 6 COURSE BEARING PONTER .............................................................................................................................. 6 TO/FROM FLAG INDICATOR ............................................................................................................................. 6 DUAL GLIDE SLOPE POINTERS ........................................................................................................................ 6 GLIDE SLOPE SCALES ........................................................................................................................................ 6 COURSE DEVIATION BAR ................................................................................................................................. 6 COURSE DEVIATION SCALE ............................................................................................................................. 6 HEADING BUG...................................................................................................................................................... 6 SLAVING AND ACCESSORY AND COMPENSATOR UNIT ............................................................................... 7 MANUAL/AUTOMATIC (FREE/SLAVE) COMPASS SLAVE SWITCH .......................................................... 7 CW/CCW COMPASS MANUAL SLAVE SWITCH ............................................................................................. 7 SLAVING METER ................................................................................................................................................. 7 WEATHER RADAR ................................................................................................................................................... 8 OPERATING CONTROLS AND INDICATORS .................................................................................................. 8

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN TABLE OF CONTENT SAFETY PRECAUTIONS ...................................................................................................................................... 9 MARGIN OF AVOIDANCE .................................................................................................................................. 9 EXTENDING LIFE OF THE MAGNETRON ........................................................................................................ 9 TURN-OFF PROCEDURE ..................................................................................................................................... 9 LIMITATIONS .............................................................................................................................................................. 1 AIRSPEED .................................................................................................................................................................. 1 ENGINE START CYCLE ........................................................................................................................................... 1 POWER PLANT (PT6A-114A) .................................................................................................................................. 1 OPERATING ALTITUDES ........................................................................................................................................ 2 WEIGHTS ................................................................................................................................................................... 2 FLIGHT LOAD FACTOR LIMITS ............................................................................................................................ 3 FUEL LIMITATIONS................................................................................................................................................. 3 MAX FULL RUDDER SIDESLIP DURATION .................................................................................................... 3 USE OF AVGAS ..................................................................................................................................................... 3 MINIMUM OIL QUANTITY ..................................................................................................................................... 3

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AIRFRAME

AIRFRAME GENERAL The Cessna C208B Grand Caravan is an all-metal, high wing, single-engine aeroplane equipped with a tricycle landing gear. A composite cargo pod is optional equipment.

FUSELAGE The fuselage length is 37 feet 7 inches (C208) and 41 feet 7 inches (C208B). The construction of the fuselage is a conventional formed sheet metal bulkhead, stringer and skin referred to as semi monocoque. Stringers are long metal strips which run the length of the fuselage and are fastened to the rings and bulkheads. The principle of the monocoque design is that the skin is designed to carry the loads and stresses. The circular shape is being maintained by the use of frames and stringers. In addition, the stringers are designed to stiffen the skin. Longerons take up the end loadings due to bending.

Major components of the structure are: •

THE FRONT AND REAR CARRY-THROUGH SPAR AND BULKHEAD. The front carry-through spar and bulkhead is an integral fail-safe structure with forgings at the top for attaching the front wing spar and forgings at the bottom for the attaching the wing strut.

•

THE REAR CARRY-THROUGH SPAR AND LANDING GEAR BULKHEAD The rear carry-through spar and landing gear bulkhead is an integral fail-safe structure with forgings at the top for attaching the rear wing spar and forgings at the bottom for attaching the main landing gear trunnions.

•

THE FORWARD DOOR POSTS The forward door post provides the load path for transferring the loads from the engine mount directly to the primary structure.

Fail-safe construction assures that the structure is designed and built in such a way that should any structural component fail, the remaining structure is capable of carrying certified limit flight loads.

WINGS The wing span is 52 feet 1 inch. The externally braced wings are constructed of a front and rear spars, formed sheet metal ribs, doublers and stringers. Ribs maintain the aerodynamic shape of the wing and transfer loads to the skin and spars. The entire structure is covered with aluminium skin.

The primary wing spars, wing carry-through spars in the fuselage, wing struts and attaching structure are of fail-safe construction for limit flight loads. The front spar is equipped with wing-to-fuselage and wing-to-strut attach fittings. The rear spar is equipped with wing-to-fuselage attach fittings.

Both wings contain an integral fuel tank. These tanks are formed by the front and rear spars, upper and lower skins and inboard and outboard closeout ribs. The inboard tank ends are located 18 inches from the wing root to form a “dry bay” for keeping fuel away from the cabin in an accident. The wing is designed to accept impact outboard of the fuel tank while minimising damage to the fuel tank area.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AIRFRAME EMPENNAGE The height is 14 feet 10 inches. 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 horizontal stabiliser is constructed of a forward and aft spar, ribs and stiffeners four upper and four lower skin panels and two left and right wraps around skin panels which also form the leading edges.

A problem of marginal nose-down elevator power was observed in transitional out-of-trim flight evaluations. This was alleviated by a single row of vortex generators fitted on top of the horizontal stabiliser just forward of the elevators. They enhance nose down elevator and trim authority. The generators are intended to prevent flow separation over the elevators at slow speeds.

CARGO POD The pod attaches to the bottom of the fuselage with screws and can be removed. The pod is fabricated with a Nomex inner housing, a layer of Kevlar and a fibreglass outer layer.

The pod has a load-carrying capacity of 1090 pounds. It has 4 separate compartments divided by aluminium bulkheads. Each compartment has a maximum floor loading of 30 lbs/sq. Ft. The maximum weight for each compartment is as follows : Forward compartment

230 lbs.

Centre compartment fwd.

310 lbs.

Centre compartment aft

270 lbs.

Aft compartment

280 lbs.

Each compartment has an individual loading door. Each door is secured in the closed position by 2 handles which latch the doors when rotated 90° to the horizontal position

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN FLIGHT CONTROLS

FLIGHT CONTROLS The flight controls consist of a conventional aileron, elevator and rudder control surfaces and a pair of spoilers mounted above the outboard ends of the flaps. The flight control surfaces are manually operated through mechanical linkages. All control cables are of stainless-steel construction.

AILERONS Round-nose ailerons are of conventional formed sheet metal ribs and smooth aluminium skin construction. 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. The servo tab on the left aileron provides reduced manoeuvring control wheel forces. Fences inboard on the ailerons on the model C208 and C208A only, enhance lateral stability.

SPOILERS Due to the long wing and greater fuel load further out in the wings, spoilers are fitted to improve lateral control at low speeds by disrupting lift over the appropriate flap. The effect is more enhanced in a flap down configuration. This feature, in turn, permits shorter length ailerons to accommodate the extremely long-span wing flaps needed for meeting the FAA’s maximum stall speed limit of 61 knots.

The spoilers are interconnected to the aileron system through a push-rod mounted on an arm on the aileron. Movement for the first 5° of aileron travel is negligible. Once the aileron has been deflected upward past the 5°-point, spoiler travel is proportional to aileron travel. The spoiler is fully retracted when the aileron is deflected downward. Maximum deflection angle is 40°.

ELEVATORS 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. Both elevator tip leading edge extensions provide aerodynamic balance and incorporate balance weights.

Elevator trimming is accomplished through two elevator trim tabs attached to the trailing edge of each elevator by full length piano-type hinges. To minimise the effects of the strong slipstream, each trim tab is located as far outboard on the elevator as practical. Dual pushrods from each actuator transmit actuator movement to dual horns on each elevator trim tab to provide tab movement. The elevator trim tabs are connected to a vertically mounted trim wheel on the control pedestal.

Trimming is manually or electrically. The left half of the switch provides power to engage the trim servo clutch. The right half controls the direction of motion of the trim servomotor. Operation of the electric trim will disconnect the autopilot.

RUDDER 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.

Trimming is accomplished through the nose wheel steering bungee mechanically connected to the rudder control system and a horizontal trim control wheel mounted on the control pedestal. The trimming system acts against the steering bungee to displace the pedals and move the rudder.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN FLIGHT CONTROLS CONTROL LOCKS A control lock secures the ailerons in the neutral position and the elevators in slightly trailing-edgedown position. Ensure that the flag is over the sidewall switch panel.

A rudder lock locks the rudder in the neutral position. Earlier serial models are equipped with a rudder lock which is operated by a spring-loaded, pull-type T-handle located on the bottom of the instrument panel to the right of the control pedestal. An interlock between the rudder lock and the fuel condition lever prevents locking the rudder when the fuel condition lever is in any position other than CUTOFF. The handle is released from the locked position by rotating it 90° and allowing it to retract forward to the unlocked position.

Later serial models incorporate an external rudder gust lock located on the left side of the tail cone. A fail-safe connection automatically disengages the lock when the elevator is deflected upward about one-fourth of its travel from neutral. Because of the fail-safe system, the elevator lock should always be engaged prior to engaging the rudder lock when securing the aircraft after shutdown. The rudder lock should be disengaged before towing, starting the engine or moving the aircraft on the ground in any manner.

WING FLAP SYSTEM The wing flaps are large span, single-slotted, semi-fowler type flaps. The flaps span 70% of the wing, increasing the area by 15%. The outboard portions have rubber leading edge vortex generators and a trailing edge lip.

The flaps are driven by an electric motor. The wing flaps are extended or retracted by positioning the WING FLAP SELECTOR LEVER on the control pedestal to the desired position. A slotted panel provides mechanical stops at 10° and 20° positions. A white tipped pointer provides flap position indication.

The system is protected by the FLAP MOTOR circuit breaker on the circuit breaker panel.

STANDBY FLAP SYSTEM A standby system can be used to operate the flaps if the primary system has a malfunction. It consists of the following: a standby flap motor, a guarded standby flap motor switch and standby flap motor up/down switch. The guarded NORM position permits operation of the flaps using the selector on the control pedestal. The STBY position disables the dynamic breaking of the primary flap motor when the standby flap motor system is operated. The UP/DOWN switch has a UP, centre-off and DOWN position.

To operate the flaps with the standby system : • • •

Select the flap lever to the desired position.

Lift the guard and place the standby flap motor switch in the STBY position.

Actuate the standby flap motor UP/DOWN switch momentarily to UP or DOWN as desired.

Observe the flap position indicator to obtain the desired position. As the standby flap system has no limit switches, actuation of the UP/DOWN switch should be terminated before the flaps reach full up or down travel, otherwise damage to the standby flap motor mounts may result.

Use of the standby flap system should be avoided with the autopilot engaged since it will cause the trim to run in opposite direction to the autopilot inputs.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN COCKPIT ARRANGEMENT

COCKPIT ARRANGEMENT INSTRUMENT PANEL The flight instruments layout is designed around the basic “T” configuration. Immediately to the left of the flight instruments are the clock, propeller anti-ice ammeter, suction gage, volt/ammeter, volt/ammeter selector switch, propeller overspeed governor test switch, left air vent pull knob, left air vent outlet and microphone and headset jacks.

The lower left side of the instrument panel contains a switch panel for the switches necessary to operate the aircraft systems. These include toggle switches for the exterior and interior lights, de-ice and anti-ice systems.

Below the flight instrument panel are the parking brake, lighting switch panel and inertial separator control.

Above the flight instrument panel are the annunciator panel, annunciator panel day/night switch, annunciator test switch and fire detector test switch.

Avionics equipment is placed vertically in dual stacks approximately in the center and just to the right of the instrument panel.

Located above the avionics stacks in the top center of the instrument panel are the engine instruments consisting of the torque indicator, propeller RPM indicator, ITT indicator, Ng% RPM indicator, oil pressure/oil temperature gage, fuel flow indicator and left and right fuel quantity indicators.

Directly to the right of the right flight instrument panel are the hour meter, right fresh air vent pull knob, right fresh air vent outlet and microphone and headset jacks and map compartment.

Located below the avionics stacks are the cabin heat switch and control panel. Provisions for air conditioning controls are also provided.

CONTROL PEDESTAL A control pedestal extending from the center of the instrument panel to the floor contains the emergency power lever, power lever, propeller control lever, fuel conditioning lever and wing flap selector and position indicator. A quadrant friction lock is located on the right side of the pedestal.

The elevator trim control and position indicator is located on the left side of the pedestal. The aileron and rudder trim controls and position indicators, the fuel shutoff valve control and cabin heat firewall shutoff valve control and a microphone are located on the lower end of the pedestal.

LEFT SIDEWALL SWITCH AND CIRCUIT BREAKER PANEL Control switches for the engine, electrical system including circuit breakers and avionics system are located on a separate panel mounted on the left cabin sidewall adjacent to the pilot.

OVERHEAD PANEL The left and right fuel tank selectors, standby flap motor switches, left and right vent air control knobs, left and right vent air outlets are located on the overhead panel. Provisions for an oxygen shutoff valve pressure gage are also provided.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN COCKPIT ARRANGEMENT PILOTS SEATS The six-way adjustable pilot’s seats may be moved forward or aft, adjusted for height and the seat angle changed. Position the seat by pulling on the small T-handle under the center of the seat bottom and slide the seat into position; then release the handle and check that the seat is locked into place.

Raise or lower the seat by rotating a large crank under the front right corner of the seat. Seat back angle is adjusted by rotating a small crank under the front left corner of the seat. The seat bottom angle will change as the seat back angle changes.

Later serial aircraft seats are equipped with armrests which can be moved to the side and raised to a position beside the seat back for stowage.

Both pilots’ seats are equipped with a five-point restraint system which combines the function of conventional type seat belts, a crotch strap and an inertia reel equipped double-strap shoulder harness in a single assembly.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN CABIN DOORS

CABIN DOORS CABIN ENTRY DOORS Entry to and exit from the aircraft is accomplished through an entry door on each side of the cabin and on the Passenger version only, through a two-piece, airstair type on the right side. A cargo door on the left side can also be used for cabin entry.

CREW ENTRY DOORS The left crew entry door incorporates a conventional interior and exterior door handle, a lock override knobs a key-operated door lock and an openable window. The right crew entry door incorporates a conventional interior and exterior door handle and manually operated inside door lock.

To open either door from the outside, rotate the handle down and forward to the OPEN position. To close the door from the inside, place the handle in the CLOSE position and pull the door shut; then rotate the handle forward to the LATCHED position. When the handle is rotated to the LATCHED position, an over-center action will hold it in position.

The lock override knob on the inside of the left door provides a means of overriding the outside door lock from inside the aeroplane. To operate the override, pull the knob and rotate it in the placard direction to unlock or lock the door.

PASSENGER ENTRY DOOR The passenger door consists of an upper and lower section. When opened, the upper section swings upward and the lower section drops down providing integral steps. The upper door section incorporates a conventional exterior door handle with a separate key operated lock, a push button exterior door handle release and an interior door handle which snaps into a locking receptacle. If the upper door is not properly latched, a red light labelled DOOR WARNING will illuminate on the annunciator panel.

The lower door features a flush handle which is accessible from either inside or outside. This handle is designed so that when the upper door is closed, the handle cannot be rotated to the open position.

The lower door also contains integral door support cables and a door lowering device.

To enter the aircraft through the passenger entry door, depress the exterior pushbutton door release, rotate the exterior door handle on the upper door section clockwise to the open position and raise the door to the over-center position. Release the lower section by pulling up on the inside door handle and rotating the handle to the open position. Lower the door section until it is supported by the integral support cables.

Do not use the outside key lock to lock the door prior to flight since the door could not be opened from the inside if it were needed in an emergency evacuation.

CARGO DOORS A two-piece cargo door is located on the left side of the fuselage. The door is divided in an upper and a lower section. When opened, the upper section swings upwards and the lower section swings forward. The upper section of the door incorporates a conventional exterior door handle with a separate key operated lock and on the Passenger version only, a pushbutton exterior emergency door handle release and an interior door handle which snaps into a locking receptacle. If the upper door is not properly latched, a red light labelled DOOR WARNING, will illuminate on the annunciator panel.

The lower door features a flush handle which is accessible from either inside or outside. This handle is designed so that when the upper door is closed, the handle cannot be rotated to the open position.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN CABIN DOORS In an emergency do not attempt to exit the Cargo version through the cargo doors without outside assistance. Since the inside of the upper door has no handle, exit from the aeroplane through these doors without outside assistance is not possible.

Do not use the outside key lock to lock the door prior to flight since the door could not be opened from the inside if it were needed in an emergency evacuation.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN CABIN DOORS

INTENTIONALLY LEFT BLANK

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN WARNING SYSTEM

WARNING SYSTEMS ANNUNCIATOR PANEL A green coloured lamp indicates a normal or safe condition in the system. An amber lamp indicates a cautionary condition exists which may or may not require immediate corrective action. A red lamp indicates a hazardous condition requiring immediate corrective action.

Two annunciator panel function switches, labelled LAMP TEST and DAY/NIGHT are located to the left of the panel. By pressing LAMP TEST all annunciators are illuminated and both fuel-selector-off warning horns are activated. With the DAY/NIGHT SWITCH in the DAY position, any annunciator that is illuminated will be at full intensity. In the NIGHT position it provides intensity down to a pre-set minimum dim level for the green lamps and the following amber lamps: • • •

Left and right fuel low.

Standby electrical power inoperative.

Standby electrical power on.

The intensity can be controlled by the ENG INST lighting rheostat. A lamp may be extinguished by pushing on the face of the light assembly and allowing it to pop out. To reactivate the annunciator, pull the light assembly out slightly and push it back in.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN WARNING SYSTEM

ANNUNCIATOR

CAUSE FOR ILLUMINATION

ENGINE FIRE

An excessive temperature condition and/or possible fire has occurred in the engine compartment.

OIL PRESS LOW GENERATOR OFF EMERGENCY POWER LEVER AUX FUEL PUMP ON

Engin oil pressure at 40 ±20 psi.

The generator is not connected to the airplane bus.

The emergency power lever is advanced out of the NORMAL position.

The auxiliary fuel pump is operating

FUEL PRESS LOW

Fuel pressure in the fuel manifold assembly is below 4,75 psi.

STARTER ENERGIZED

The starter-generator is operating in starter mode.

IGNITION ON

Electrical power is being supplied to the engine ignition system.

VOLTAGE LOW

Electrical system bus voltage is below 24,5 volts, and power is being supplied from the battery.

VACUUM LOW

The vacuum system suction is less than approximately 3.0 in Hg.

RESERVOIR FUEL LOW

The fuel level in the reservoir tank is approximately one-half full or less.

LEFT FUEL LOW

Fuel quantity in the fuel left tank is 25 gallons or less.

RIGHT FUEL LOW

Fuel quantity in the right left tank is 25 gallons or less.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN WARNING SYSTEM

ANNUNCIATOR STBY ELEC POWER INOP INVERTER INOP

CAUSE FOR ILLUMINATION Electrical power is not available from the standby alternator.

The selected inverter has failed or has not been turned on (if installed).

FUEL SELECT OFF

One or both fuel tank selectors are off, the fuel selector warning circuit breaker is not set or the start control circuit breaker is not set.

DOOR WARNING

Indicates the upper cargo door and/or upper aft passenger door (Passenger Version Only) are not latched.

BATTERY OVERHEAT

The electrolyte temperature in the nicad battery is critically high (1600 F).

BATTERY HOT

The electrolyte temperature in the nicad battery is high (1400 F).

CHIP DETECTOR

Indicates that metal chips have been detected in either the reduction gearbox case or accessory gearbox case (if installed).

STBY ELEC POWER ON

The standby alternator is supplying electrical power to the bus.

WINDSHIELD ANTI – ICE DE – ICE PRESSURE A/P OFF A/P OFF

A/P OFF A/P BC

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Electrical power is being supplied to the windshield anti-ice power relay.

Pressure in the device boot system has reached approximately 15 psig.

The autopilot has disengaged.

The autopilot is in back course mode.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN WARNING SYSTEM FUEL SELECTORS OF WARNING SYSTEM The system consists of redundant warning horns, a red annunciator light labelled FUEL SELECT OFF, actuation switches and electrical hardware. The dual aural warning system is powered through the START CONT circuit breaker. A non-pullable FUEL SEL WARN circuit breaker is installed in series to protect the integrity of the starter system. The annunciator is powered from the ANNUN PANEL circuit breaker. The warning system functions as follows : •

ABOTH SELECTORS OFF The red FUEL SELECT OFF annunciator illuminates. One fuel selects off warning horn is activated.

•

DURING ENGINE START WITH EITHER SELECTOR OFF The red FUEL SELECT OFF annunciator illuminates. Both fuels select off warning horns are activated.

•

ONE SELECTOR OFF AND FUEL REMAINING IN THE TANK BEING USED IS LESS THAN 25 GALLONS The red FUEL SELECT OFF annunciator illuminates. One fuel selects off warning horn is activated.

If the FUEL SEL WARN circuit breaker has popped or the START CONT circuit breaker has been pulled, the red FUEL SELECT OFF annunciator will illuminate even with both tanks ON.

STALL WARNING SYSTEM A vane-type stall warning unit is fitted in the leading edge of the left wing. The unit is electrically connected to the stall warning horn located overhead of the pilot's position. It senses a change in airflow over the wing and operates the warning horn at airspeeds between 5 and 10 knots above the stall in all configurations. The system is protected by the STALL WRN circuit breaker. The vane and sensor unit is equipped with a heating element operated by the STALL HEAT switch and is protected by the STALL WRN circuit breaker.

OVERSPEED WARNING SYSTEM An airspeed pressure switch in the pitot-static system is used to actuate an airspeed warning horn in the event of excessive airspeed. The horn is located behind the headliner in the area above the pilot and will sound when the airspeed exceeds Vmo (175 KIAS). A warning signal will also be heard in the pilot’s headsets.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN WARNING SYSTEM

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN LANDING GEAR AND BRAKE SYSTEM

LANDING GEAR AND BRAKE SYSTEM LANDING GEAR The landing gear is a fixed-gear, tricycle type with steerable nose wheel and two main wheels. To improve operation from unpaved surfaces, the standard nose gear fork and tyres can be replaced by a 3" extended nose gear and oversized tyres. The main landing gear has a “tear away” feature which reduces damage to the fuselage structure.

Main gear shock absorption is provided by tubular spring-steel main landing gear struts and an interconnecting spring-steel tube between the two main landing gear struts. The interconnecting tube reduces bending in the fuselage landing gear bulkheads.

Nose gear shock absorption is provided by an oil-filled shock strut combined with a drag link spring providing vertical and aft displacement restraint. Fail-safe features of the drag link include absorption of some energy even with complete oil loss and retains the nose gear in the event of torque link failure.

Nose gear steering is accomplished by using the rudder pedals. A spring-loaded steering bungee, which is connected to the nose gear and to the rudder bars, will turn the nose gear through an angle of 15° each side of centre. By applying differential breaking the degree of turn may be increased up to 56° each side of centre. If the nose wheel is turned beyond its limits during towing, a frangible stop will fracture and the over travel indicator block hanging on a small cable, will fall into view alongside the nose strut.

BRAKE SYSTEM A single-disc, hydraulically-actuated brake is fitted on the inboard side of each main landing gear wheel. Each brake is connected by a hydraulic line to a master cylinder attached to each pilot's rudder pedal. To set the park brake, apply brake pressure and pull the handle aft. To release the parking brake, push the handle fully in.

A brake fluid reservoir located just forward of the firewall on the left side of the engine compartment provides additional brake fluid for the brake master cylinders. The fluid in the reservoir should be maintained between the MIN and MAX level prior to each flight. The brake system uses MIL H-5606 hydraulic fluid. This fluid is a mineral base type and is a flammable petroleum product and coloured red. Systems using mineral base fluids incorporate synthetic rubber seals.

Later model aeroplanes (208B01030 and above) have metallic type brakes. When conditions permit, hard braking is beneficial in that the resulting higher brake temperatures tend to maintain proper brake glazing and will prolong the expected brake life. The habitual use of light and conservative brake application is detrimental to metallic brakes.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN LANDING GEAR AND BRAKE SYSTEM

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN TURBINE ENGINE THEORY

TURBINE ENGINE THEORY

Turbine engines are made up of the following sections: • • • • • • •

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Air inlet.

Compressor.

Combustion section.

Turbine.

Exhaust.

Accessory section.

Support systems (starter, lubrication, fuel and auxiliary components such as anti-icing, airconditioning and pressurisation).

TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN TURBINE ENGINE THEORY AIR INLET DUCT The purpose of the air inlet duct is to channel the incoming air to the compressor with the least possible loss of energy. In addition to the design of the inlet, three other variables determine how much air will pass through the compressor: • • •

Ambient air density.

Airspeed of the aircraft.

Rotational speed of the compressor.

COMPRESSORS Two main functions of the compressor are to supply compressed air to the burners plus supply bleed air to fulfil various engine and airframe system requirements. To supply compressed air to the burners, the compressor increases the pressure of the mass of air that is channelled through the air inlet ducts and routes it to the burners in the amount and pressure required.

Bleed air might be tapped off any of the various compressor stages. With each successive stage the pressure and temperature of the air increases. The specific stage for bleed air extraction is determined by the temperature and pressure required for the job.

THE CENTRIFUGAL COMPRESSOR The compressor accelerates the airflow by slinging it outward from the centre of the compressor wheel. The primary components of the centrifugal compressor are the impeller (rotor), a diffuser (stator) and a compressor manifold. The impeller catches the incoming air and accelerates it outward toward the diffuser. The purpose of the diffuser vanes is to direct the air from the impeller to the manifold at the gentlest angle possible to retain the maximum amount of energy. Another purpose of the diffuser is to provide the combustion chamber with air at the correct velocity and pressure for maximum efficiency. The air is allowed to expand into a divergent duct. As the air spreads out, the velocity drops and the static pressure increases. Multiple stages are connected by a system of convergent and divergent ducts to compress and direct the airflow to succeeding stages at the proper angle.

Advantages of centrifugal compressors : • • • • • •

High pressure rises per stage. Up to 10:1 in a single stage.

Good efficiency (compression) over a wide rotational speed range, idle to full power.

Simplicity of manufacture and low cost.

Low weight.

Low starting power requirements.

More resistant to foreign object damage.

Disadvantages of centrifugal compressors : • •

Large frontal area for a given airflow.

More than two stages are not practical due to energy loss in the airflow when making the turn from one impeller to the next.

THE AXIAL FLOW COMPRESSOR The airflow follows a path parallel to the axis of rotation, hence the name. Airflow is compressed by accelerating it across multiple compressor stages. Airflow passes through a convergent duct to force it into an even smaller area. The size of the compressor blades decreases from the first to the last to follow the narrowing design of the convergent duct.

The main elements of an axial-flow compressor are rotors and stators. Rotor blades are fixed on a rotating spindle and force air rearward in the same manner as a fan. The rotor blades are pre-set at a ISSUE 01 – REV. 00

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN TURBINE ENGINE THEORY specific angle and are contoured similar to small propeller blades. Stators, which are stationary bladetype air foils, are located between each compressor stage. The stators act as a diffuser in each stage converting part of the energy air into pressure by slightly slowing down the air. They also direct the airflow into succeeding compressor stages at the correct angle.

The first set of stators, located before the compressor stage, is referred to as inlet guide vanes. A similar set of guide vanes, called straightening vanes, might be located at the compressor exit to stop the exiting air mass from rotating as it moves forward into the combustion chamber.

Advantages of axial flow compressors : • • •

High peak efficiencies (compressor pressures ratio) created by its straight through design allowing higher ram efficiency.

Higher peak efficiencies (pressure) attainable by addition of compression stages if desired.

Small frontal area for a given airflow resulting in low drag.

Disadvantages of axial flow compressors : • • • • • •

Difficulty and high cost of manufacture.

Relatively high weight.

High starting power requirements.

Low pressure rises per stage.

Good efficiency only over a narrow rotational speed range.

Very susceptible to foreign object damage.

COMPRESSOR STALLS The angle at which the incoming air meets the compressor blades varies as a result of the air’s velocity and the compressor’s rotational speed. The blade’s combined influence forms a vector that is called the angle of attack. An imbalance between the two vector quantities might stall the airflow.

By definition, a compressor stall occurs when the air mass travelling through the compressor slows down and stops. In extreme cases the airflow might even reverse direction. An engine compressor stall may be noted by a single or multiple loud “popping” noise from the engine compartment. The situation may be resolved by reducing the power to a point where the “popping” discontinues and slowly advancing the throttle to the required setting. The use of bleed air may also help eliminate engine compressor stalls if this situation is encountered.

Compressor blades stall for numerous reasons : • • • • •

Anything that alters the proper operation of the compressor blades such as blade failure, foreign object damage, etc.

A fuel mixture that is to lean.

Abrupt pitching movement.

Excess fuel flow.

Engine operating speeds well above or below normal recommended operating speed.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN TURBINE ENGINE THEORY THE COMBUSTION CHAMBER The combustion chamber houses the fuel nozzles and igniters. Fuel is mixed with the compressed air in the combustion chamber. The temperature of the air as it leaves the compressor is ±315°C. After being processed by the combustion chamber, the air temperature is ±870°C. To accomplish its task of efficiently burning the fuel/air mixture, the combustion chamber must: • • •

Mix the fuel and air in the proper manner for the ambient conditions to assure the best possible combustion.

Cool the hot gasses to a temperature within the normal operating range of the turbine wheel.

Channel the exhaust gases in such a manner that it maximises turbine wheel effectiveness.

The combustor lining is designed for proper flow of air into and out of the combustor which provides for proper flame patterns. No flame is allowed to touch the combustor lining. In addition, the combustor lining assures that only a small amount of the total airflow (approximately 25%) is mixed with fuel and burned. The rest of the airflow provides cooling and dilution of the products of combustion.

TURBINE A turbine section is designed to extract as much energy as possible from the high velocity airflow issuing from the combustor. The sole purpose is to convert the kinetic energy of combustion chamber exhaust gasses into mechanical energy to operate the compressor and other accessories.

Approximately 60 - 80 percent of the total energy of the exhaust gasses go to that purpose. The turbine section is a divergent duct. Each stage is larger in diameter than the preceding stage to compensate for loss of energy through each stage. This allows an equal sharing of the load between stages.

Stators or turbine guide vanes are located in front of and between each turbine stage to smooth the airflow and properly direct it to the next turbine wheel. High energy airflow strikes the turbine blades at an angle. As the airflow does not easily change direction, it imparts a pressure upon the turbine blades which causes a turning force upon the wheel. The turbine stages are mounted on the same shaft with the compressor stages.

THE EXHAUST The exhaust section, which is located directly behind the turbine section has two primary purposes: produce exhaust gas high exit velocity and minimise exhaust gas turbulence. The purpose of the exhaust cone is to channel all exhaust gasses as they are discharged from the turbine blades and combine them into a single, cohesive gas stream. Fixed struts also serve as a stabilising influence on the gas and minimise the swirling motion. If this were not done, exhaust gasses would exit the engine at an approximate 45° angle resulting in unwanted drag.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN TURBINE ENGINE THEORY TURBOPROP ENGINE POWER CONVERSION There are two power output turbine designs: the fixed shaft design and the free turbine engine.

NEW

THE FIXED SHAFT DESIGN

One central shaft mounts the compressor and the turbine. The shaft extends forward into the gearbox where high RPM is converted to low RPM suitable to drive the propeller. All rotating components within a fixed shaft engine rotate together. Advantages of a fixed shaft system include controlled descent capability due to prevention of windmilling overspeed and rapid reverse thrust availability.

NEW

THE FREE TURBINE

An airflow couple is utilised between the gas generator and power turbine shaft. The gas generator turbine powers the compressor. The power turbine drives into the reduction gearbox in the same manner as the fixed turbine. The gearbox converts high RPM low torque input into usable low RPM high torque power. Because a free turbine system allows the propeller to operate independent of the compressor, there are several advantages : • • • • •

Better control of propeller speed.

The propeller can be held at very low rpm during taxiing, with low noise and low blade erosion.

The engine is easier to start, especially in cold weather.

The propeller and the gearbox do not directly transmit vibrations into the gas generator.

A propeller brake can be installed during turnarounds.

PT6 OPERATING PRINCIPLES Air enters the engine through an annular plenum chamber formed by the compressor inlet case from where it is directed to the compressor. The compressed air passes through diffuser ducts which turn it 90° in direction. It is routed through straightening vanes to the combustion chamber. The air changes direction to enter the combustion chamber liner where it reverses direction and is mixed with fuel.

Fuel is injected into the combustion chamber via 14 simplex (single orifice) nozzles. The mixture is then ignited by two igniters. Expanding gases reverse direction and pass through the compressor turbine guide vanes to the compressor turbine. The still expanding air passes through a second set of stators to drive the power turbine. Gas from the power turbine is exhausted via the exhaust.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT

POWER PLANT NEW DESCRIPTION The Pratt & Whitney Canada Inc. or PT6A-114A (C208B) is a free turbine engine with two independent turbines. One driving a compressor in the gas generator section. A second driving the reduction gearing for the propeller. The compressor section has 3 axial stages and 1 centrifugal stage. One stage compressor turbine. One stage power turbine.

The engine is flat rated at 675 SHP (1865 foot-pounds of torque at 1900 RPM varying linearly to 1970 foot-pounds of torque at 1800 rpm. Below 1800 RPM, the maximum torque remains constant at 1970 foot-pounds). Torque is the measured amount of shaft horsepower delivered to the propeller. The torque bending force is caused by the resistance of the air pushing against the advancing blade. At a lower rpm the blade angle is increased to take a bigger “bite” of air therefore increasing the resistance and hence the torque.

The gas generator speed is 37,500 rpm at 100% Ng. Max permissible speed is 38,100 rpm at 101.6% Ng. The power turbine speed is 33,000 RPM at a propeller shaft speed of 1900 RPM. The engine is supported in a steel 9-element space frame. This frame attaches to the firewall at 5 points. To minimise directional and longitudinal deviations with changes in thrust, the engine is canted at 3° nose-down and 5° to the right.

ENGINE CONTROLS POWER LEVER The power lever is connected through linkage to a cam assembly mounted in front of the FCU. It controls engine power, via pneumatic control of the metering valve, from maximum take off power back through idle to full reverse thrust. The power lever has MAX, IDLE, BETA and REVERSE range positions. The lever selects propeller pitch when in BETA range. The BETA range enables the pilot to control propeller blade pitch from idle thrust to through a zero or no thrust condition to maximum reverse thrust.

Do not move the power lever into the BETA range when the propeller is feathered as the propeller reversing linkage can be damaged.

EMERGENCY POWER LEVER The emergency power lever is connected through linkage to the manual override lever on the FCU. It governs fuel supply to the engine should a pneumatic failure in the FCU occur. A pneumatic failure will result in the fuel flow decreasing to idle (48% Ng at sea level and 52% Ng at 5000 ft).

The emergency power lever has NORMAL, IDLE and MAX positions. The NORMAL position is used for all normal engine operation when the FCU is operating normally and power is selected by the power lever. The range from IDLE to MAX positions governs engine power and is used when a pneumatic malfunction has occurred in the FCU 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 position.

The knob has cross hatching and is visible when the lever is in the MAX position. Also, the emergency power lever is enunciated on the annunciator panel when the lever is out of the normal position. These precautions are intended to preclude starting of the engine with the emergency power lever in any position other than NORMAL. The emergency power lever and its associated manual override system is considered to be an emergency system and should only be used in the event of a fuel control pneumatic malfunction.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT For starting ensure that the emergency power lever is in the normal position otherwise an over temperature or hot start may result.

When using this lever in the override position engine response may be more rapid than when using the power lever as the fuel is not metered. Utilise slow and smooth movement of the emergency power lever to avoid engine surges and/or exceeding ITT, Ng and torque limits. The emergency power lever may have a dead band such that no engine response is observed during the initial forward travel from the IDLE position. When using the emergency power lever, monitor gas generator RPM when reducing power near idle, to keep it from decreasing below 65% Ng in flight.

PROPELLER CONTROL LEVER The propeller control lever is connected to the propeller control governor mounted on the front section of the engine. It controls propeller governor settings from maximum RPM 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 desired propeller RPM for the cruise. The FEATHER position is used for shutdown to stop rotation of the power turbine and the front section of the engine. Once the gas generator is shut down no lubrication is available and rotation of the forward section 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.

The propeller lever operates a speeder spring inside the primary governor to reposition the pilot valve, which results in an increase or decrease of propeller rpm. For propeller feathering, the propeller control lever lifts the pilot valve to a position that causes complete dumping of high-pressure oil, allowing the counterweights and feathering spring to change the pitch.

FUEL CONDITION LEVER The fuel condition lever is connected through linkage to a combined lever and stop mechanism on the FCU. The lever and stop also function as an idle stop for the FCU rod. It 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 cuts all fuel to the fuel nozzles. LOW IDLE positions the control rod stop to provide an RPM of 52% Ng. HIGH IDLE provides an RPM of 65% Ng. QUADRANT FRICTION LOCK A quadrant friction lock, located on the right side of the pedestal, is provided to minimise creeping of the engine controls once they have been set. The lock increases the friction on the engine controls when rotated clockwise.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT ENGINE INSTRUMENTS TORQUE INDICATOR The torque indicator indicates the torque being produced by the engine. One some Cargo versions the torque-indicator system is powered by 28-volt DC power. On the other cargo versions and passenger versions, the torque indicator is pressure actuated.

Two independent lines enter the back of the torque indicator. One line measures engine torque pressure and the other the reduction gearbox internal pressure. The torque indicator monitors the engine torque pressure and converts this pressure into an indication of torque in ft-lbs.

Normal operating range (green arc) is from 0 to 1865 ft-lbs. The alternate power range (striped green arc) is from 1865 ft-lbs to 1970 ft-lbs. The maximum torque (red line) is 1970 ft-lbs and maximum take-off torque (red wedge) is denoted by "T. O" at 1865 ft-lbs.

PROPELLER RPM INDICATOR 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. Normal operating range (green arc) is from 1600 to 1900 rpm. The maximum (red line) is 1900 rpm.

ITT INDICATOR The ITT (inter-turbine temperature) indicator displays the gas temperature between the compressor and the power turbines. Normal range (green arc) is 100°C to 740°C. Engine operations which exceed 740°C may reduce engine life. A maximum (red line) at 805°C. Maximum starting temperature (red triangle) is 1090°C. These limitations are restricted to 2 seconds. Although not indicated, the idle ITT is restricted to 685°C. Ng% RPM INDICATOR The Ng % indicator indicates % of gas generator rpm based on a figure of 100% at 37,500 rpm.

It is electrically operated from the gas generator tachometer-generator mounted on the lower righthand side portion of the accessory case. The normal operating range (green arc) is from 52% to 101.6% Ng. The maximum (red line) is at 101.6% Ng. FUEL FLOW INDICATOR The fuel flow indicator indicates fuel flow in lbs/hr based on Jet A fuel. Fuel is measured downstream of FCU just before it enters the flow divider. When the power is removed from the indicator, the needle will stow below the zero in the OFF band. The instrument is electrically operated.

OIL PRESSURE GAGE A direct pressure oil line from the engine delivers oil at operating pressure to the indicator. Minimum pressure (red line) is 40 psi. Normal operating range (green arc) is from 40 to 85 psi. The maximum pressure (red line) is at 105 psi.

Normal oil pressure is 85 to 105 psi at gas generator speeds above 72% with the oil temperature between 60° and 70°C. Oil pressures below 85 psi are undesirable and should be tolerated only for the completion of the flight, preferable at a reduced power setting. Oil pressures below 40 psi are unsafe and require that either the engine be shut down or a landing be made as soon as possible using minimum power as required to sustain flight.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT OIL TEMP GAGE The instrument is operated by an electrical-resistance type temperature sensor which receives power from the aeroplane electrical system. Minimum temperature (red line) is -40°C. Normal operating range (green arc) is from 10° to 99°C. Maximum operating temperature (red line) is 99°C.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT ENGINE LUBRICATION SYSTEM The lubrication system consists of a pressure system, a scavenge system and a breather system. The main components include an integral oil tank at the back of the engine, an oil pressure pump at the bottom of the tank, an external double element scavenge pump at the back of the accessory case, an internal double element scavenge pump inside the accessory gearbox, an oil-to-fuel heater on the top of the accessory case, an oil filter internally of the oil tank and an oil cooler located on the right side of the nose cowl.

Oil is drawn from the bottom of the tank through a filter screen where it passes through a pressure relief valve for regulation of the pressure. The pressure oil is then delivered from the main oil pump to the oil filter. Pressure oil is then routed to the number 1 bearing and accessory gearbox, bearings number 2, 3, 4, reduction gears, torquemeter and propeller governor. Pressure oil is also routed to the oil-fuel heater where it then returns to the tank.

After cooling and lubricating all the moving parts, oil is scavenged as follows: • • • • •

Oil from the number 1 bearing sump to accessory gearbox by gravity.

Oil from the number 2 bearing sump is pumped to the accessory gearbox by number 2 scavenge pump.

Oil from the number 3 and 4 bearing sump is pumped to the accessory gearbox by the power turbine bearing scavenge pump.

Oil from the propeller governor, front thrust bearing, reduction gear accessory drives and torquemeter is pumped to the tank by the reduction gearbox scavenge pump through a thermostatically controlled air-oil cooler and then returned to the tank.

The rear element of the internal scavenge pump scavenges oil from the accessory gearbox and routes it through the air-oil cooler where it then returns to the oil tank.

The presence of pressurised air in the bearing cavities is a result of gas path air leaking across carbon and oil seals. This air pressure assists in oil return to the tank by putting a head of pressure on the scavenge oil at the bearing sumps. 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 return lines.

The oil tank capacity is 9.5 US quarts and the total system capacity is 14. US quarts (including oil in filter, cooler and hoses). An oil dipstick indicates US quarts low when the engine is hot. Fill to within 1½ qtrs. of max hot or max cold positions. To obtain an accurate oil level reading, it is recommended the oil level be checked within 10 minutes after shutdown when oil is hot (MAX).

HOT marking) or prior to the first flight of the day (MAX COLD marking). If more than 10 minutes has elapsed and oil is still warm perform an engine dry motoring run before checking the oil level.

As the oil tank is pressurised at 3 to 6 psi, ensure that the oil dipstick is securely latched down.

Operating the engine with less than the recommended oil level and with the dipstick cap unlatched will result in excessive oil loss and eventual engine stoppage.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT IGNITION SYSTEM Turbine engine ignition systems are designed to initially ignite the fuel/air mixture, which will then sustain the combustion on its own. The ignition system consists of 2 igniters, an ignition exciter, 2 high tension leads, an ignition monitor light, an ignition switch and a starter switch.

Engine ignition is provided by 2 igniters in the engine combustion chamber. The igniters are energised by the ignition exciter mounted on the right-hand side of the engine compartment. Electrical energy from the ignition exciter is transmitted through 2 high-tension leads to the igniters in the engine. A green annunciator, labelled IGNITION ON, will illuminate when electrical power is being supplied to the igniters.

Ignition is controlled by an ignition switch and a starter switch. The ignition switch has two positions, NORMAL and ON. NORMAL POSITION Arms the system so that ignition will be obtained when the starter switch is placed in the START position. It is used for ground starts and air starts with starter assist.

ON POSITION Provides continuous ignition regardless the position of the starter switch. The use of ignition for extended periods of time will reduce ignition system component life. However, the ignition should be turned ON to provide continuous ignition under the following conditions : • • • • •

Air starts without starter assist.

Operation on water or slush covered runways.

During flight in heavy rain.

Inadvertent icing encounters until the inertial bypass separator has been in bypass for 5 minutes.

Near fuel exhaustion as indicated by the RESERVOIR FUEL LOW light.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT STARTING SYSTEM The main function of the starter switch is control of the starter for rotating the gas generator portion of the engine during start. It also provides ignition during starting. The starter/generator functions as a motor for starting. It will motor the gas generator section until a speed of 46% Ng. At 46% Ng the staring cycle will automatically be terminated. The starter switch has 3 positions, OFF, START and MOTOR. OFF POSITION The ignition system is de-energised. This is the normal position at all times except during engine start or engine clearing.

START POSITION Rotates the gas generator section of the engine. Also energises the ignition system provided the ignition switch is in the NORMAL position. When the engine has started the start switch must be placed in the OFF position to activate the generator. Starter contactor operation is indicated by the amber STARTER ENERGIZED annunciator.

MOTOR POSITION Motors the engine without having the ignition circuit energised. This position is spring loaded back to the OFF position. 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.

A minimum battery voltage of 24 volts is not always an indication that the battery is near full charge or in a good condition. This is especially true for a nicad battery, which maintains a minimum no-load voltage of 24 volts even at 50% or less charge condition. Therefore, if the gas generator acceleration is less than normally observed, return the fuel condition lever to CUTOFF and discontinue the start.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT ENGINE START CONSIDERATIONS •

If no ITT rise is observed within 10 seconds after moving the fuel condition lever to LOW IDLE, or ITT rapidly rises to 1090°C, move the fuel condition lever to CUTOFF and perform the engine clearing procedure.

•

With a cold engine or after making a battery start, which causes a high initial generator load due to battery recharging, it may be necessary to advance the power lever slightly ahead of the idle detent to maintain a minimum idle of 52% Ng. Since the generator contactor closes when the starter switch is turned OFF, anticipate the increased engine load by advancing the power lever to obtain approximately 55% Ng before turning the starter switch OFF. This prevents the initial generator load from decreasing idle rpm below the minimum 52% Ng.

•

Under hot OAT and/or high ground elevation conditions idle ITT may exceed maximum idle ITT limitations of 685°C. Increase Ng and/or reduce accessory load to maintain ITT within limits.

•

If, during the start, the starter accelerates the gas generator rapidly above 20% Ng, suspect gear train decouple. Do not continue the start. Rapid acceleration through 35% Ng suggests a start on the secondary nozzles. Anticipate a hot start.

•

After an aborted start for whatever reason, it is essential before the next start attempt to allow adequate time to drain off unburned fuel. Failure to do so could lead to a hot start, a hot streak leading to hot section damage, or torching of burning fuel from the engine exhaust on the next successful ignition.

A dry motoring, within starter limitations after confirming that all fuel drainage has stopped, ensures that no fuel is trapped before the next start.

•

If the STARTER ENERGIZED annunciator fails to go out after engine start, the generator will not function because the start contactor may be stuck closed. The battery switch should be turned off and the engine should be shut down if such an indication is observed.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT 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 the ducting and an inertial separator system before entering through a circular plenum chamber where it is directed to the compressor guide vanes. The compressor air inlet incorporates a screen which will prevent entry of large particles, but does not filter the air.

INERTIAL SEPARATOR SYSTEM An inertial separator system in the air inlet duct prevents moisture particles from entering the compressor air inlet plenum when in bypass mode. It consists of two moveable vanes and a fixed air foil which, during normal operations route the air through a gentle turn into the compressor air inlet plenum.

In the BYPASS position the vanes are positioned so that the inlet air is forced to execute a sharp turn in order to enter the inlet plenum. Any moisture particles are separated from the inlet air and discharged 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 handle is labelled BYPASS-PULL, NORMAL-PUSH. The inertial separator should be in the BYPASS position under the following conditions : • •

Running the engine on the ground or in flight in visible moisture with an OAT of 4°C or less.

Ground operation on dusty, sandy fields.

With the inertial separator in the BYPASS position and power set below the maximum torque limit, decrease the maximum cruise torque by 100 foot-pounds. Do not exceed 740°C ITT. Fuel flow for a given torque setting will be 15 pph higher.

To unlock, push slightly forward and rotate the handle 90° counter clockwise. The handle can then be pulled into the BYPASS position. Once moved into the BYPASS position, air loads on the moveable vanes hold them in position.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT EXHAUST SYSTEM The primary exhaust is attached to the right side of the engine aft of the propeller reduction gearbox.

A secondary pipe, fitted over the end of the primary exhaust, carries the gasses away from the cowling into the slipstream. The juncture of the primary exhaust pipe and the secondary exhaust duct is located directly behind the oil cooler.

Since the secondary exhaust duct is of a larger diameter then the primary exhaust pipe, a venturi effect is produced by the flow of the exhaust. This venture effect creates a suction behind the behind the oil cooler which augments the flow of cooling air through the cooler. This additional air flow improves oil cooling during ground operation of the engine.

Do not leave the power lever in BETA mode for more than 30" with a right crosswind as the pod may be damaged.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT 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 valves.

The system provides fuel flow to satisfy the speed and power demands of the engine.

OIL-FUEL HEATER Fuel is delivered from the reservoir to the oil-to-fuel heater. The oil-to-fuel heater utilises heat from the engine lubrication system to preheat the fuel. 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 oil tank.

FUEL PUMP Fuel enters the EDP fuel pump through a 74-micron filter. 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 enter the pump chamber.

The pump increases the fuel pressure and delivers the fuel to the FCU via a 10-micron filter in the pump outlet. A bypass valve in the pump body enables unfiltered, high-pressure fuel to enter the FCU in the event of this filter becoming blocked.

FUEL CONTROL UNIT The FCU consists of a fuel metering section, temperature compensating section, a compressor turbine (Ng) governor, a pneumatic computing section, a manual override system and a power turbine (Nf) governor.

The FCU determines the proper fuel schedule to provide the power required as established by the power lever. This is accomplished by controlling the speed of the compressor turbine (Ng). Engine power output is directly dependent on compressor turbine speed. The FCU governs Ng, thereby actually governing the power output of the engine. Control of Ng is accomplished by regulating the amount of fuel supplied to the combustion section of the engine.

A. Metering Section The metering valve input is supplied with fuel at pump pressure (P1). The fuel pressure immediately after the metering valve is called metered fuel (P2) which flows to the fuel divider.

The bypass valve maintains a constant fuel pressure differential (P1 - P2) across the metering valve. The metering valve consists of a contoured needle operating in a sleeve and regulates the flow of fuel by varying the orifice area. The orifice area of the metering valve is changed by the valve movement to meet specific engine requirements.

Fuel pump pressure (P1) in excess of requirements is returned to the fuel pump. A fuel cut-off valve is situated downstream of the metering valve. It provides a positive means of shutting off fuel flow to the engine. During normal operation the valve is fully open and offers no restriction of fuel flow to the divider. The valve is operated by a cut-off lever which is mechanically linked to the FCU. B. Pneumatic Computing Section The computing section consists of an acceleration bellows and a governing bellows connected to a common rod. The end of the acceleration bellows is attached to the body casting and provides an absolute pressure reference. The governor bellows is secured in the body cavity and its function is similar to that of a diaphragm. Movement of the bellows is transmitted to the metering valve via a torque tube assembly. This tube is positioned during assembly to provide a force in a

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT direction tending to close the metering valve, while the bellows act against this force to open the metering valve.

Third stage turbine discharge pressure (P3) is split up into Px and Py pressure. Px and Py vary with changing engine operating conditions as well as inlet air temperature. Py pressure is applied to the outside of the governor bellows, while Px pressure is applied to the inside of the bellows and the outside of the acceleration bellows. Any change in Py will therefore have more effect on the diaphragm than an equal change in Px pressure due to the difference in effective area. When both pressures increase simultaneously, as during acceleration, the bellows move downwards and the metering valve moves in an opening direction. When Py decreases as the desired Ng is approached, the bellows will travel to reduce the opening of the metering valve. When both pressures decrease simultaneously the bellows will move upwards and reduce the metering valve opening because Py is more effective than Px. This occurs during deceleration and moves the metering valve to the minimum flow stop.

C. Compressor Turbine (Ng) Governor The power lever incorporates a speed scheduling cam which depresses an internal rod when the power lever is advanced. The governor lever is pivoted and one end operates against an orifice to form the governor valve. The enrichment lever is also pivoted and actuates a fluted pin which operates against the enrichment hat valve. The speed scheduling cam applies tension to the governor spring which applies a force to close the governor valve. The enrichment spring between the enrichment and governor levers provides a force to open the enrichment valve. As the drive shaft rotates, it in turn rotates a table on which the governor flyweights are mounted. As the Ng increases, centrifugal loading causes the flyweights to move outwards. This in turn moves the Ng platform upwards overcoming the enrichment spring force, closing the enrichment valve and opening the governor valve.

D. Manual Override System The retaining plate and cover containing the governor bellows stop are replaced by a shaft and stop assembly. If operated, it pushes against the end of the governor bellows to open the metering valve and increase the fuel flow.

E. Power Turbine (Nf) Governor In the event of a power turbine overspeed condition, a governing orifice in the Nf governing section is opened by flyweight action of the governor. Py pressure is bled off to the atmosphere.

When this occurs, Py pressure acting on the FCU governor bellows decreases and moves the metering valve in a closing direction, thus reducing fuel flow. This in turn decreases Ng speed and consequently Nf speed.

F. Operation 1. Engine Starting The starting cycle is initiated with the Power Lever placed in the IDLE position and the Fuel Condition Lever in the CUTOFF position. The ignition and starter are switched on, and when the required Ng speed is attained, the Fuel Condition Lever is placed in the LOW IDLE position. Following ignition, the engine accelerates to idle speed. During the starting sequence, the metering valve in the FCU is in a low flow position. The flow divider schedules the metered fuel from the FCU, between the primary and secondary nozzles. Fuel is delivered to the combustion chamber through 10 primary and 4 secondary nozzles. Metered fuel is delivered initially by the primary nozzles, with the secondary nozzles cutting in above a certain value.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT As the compressor accelerates, the discharge pressure (P3) increases. This creates an increase in Px and Py pressure, which is modified P3 pressure. As Py pressure acts on a greater area, the bellows are forced down causing the main metering valve to move in an opening direction. Excess fuel supplied by the fuel pump will pass via the bypass valve back to the tank. When Ng approaches idle speed, the centrifugal force of the Ng governor flyweights begins to overcome the governor spring force and open the governor valve, bleeding off Py pressure. This creates a Px/Py differential which causes the metering valve to move in a closing direction until the required idle speed fuel flow is obtained. Any variation from the selected speed will be sensed by the Ng governor flyweights and will result in an increased or decreased weight force.

2. Acceleration As the power lever is advanced above the idling position, the speed scheduling cam is repositioned moving the cam follower lever to increase governor spring force. The governor spring then overcomes the flyweights and moves the governor lever, closing the governor valve. Px and Py immediately increase, causing the metering valve to move in an opening direction.

As Ng and consequently Nf increase, the propeller governor increases the pitch of the propeller blades to control Nf at the selected speed and applies the increased power as additional thrust. Acceleration is complete when the centrifugal force of the governor flyweight again overcomes the governor spring and opens the governor valve.

3. Governing Once the acceleration cycle has been established, any variation in engine speed from the selected speed will be sensed by the Ng governor flyweights and will result in increased or decreased weight force. This change in weight force will cause the governor valve to either open or close which will be reflected by the change in fuel flow necessary to re-establish the selected speed.

4. Altitude Compensation Altitude compensation is automatic since the acceleration bellows assembly in the FCU is evacuated and affords an absolute pressure reference. Compressor discharge air (P3) is a measurement of engine speed and density. Px is proportional to P3, so it will decrease with a decrease in air density. This is sensed by the acceleration bellows which act to reduce fuel flow on acceleration at altitude.

5. Deceleration When the Power Lever is retarded, the speed scheduling cam raised, reducing the governor spring force and allows the governor valve to move in an opening direction. The resulting drop in Py pressure moves the metering valve in a closing direction until it contacts the Wf minimum flow stop. This stop ensures sufficient metered fuel flow to the engine to prevent flameout. The engine continues to decelerate until the governor flyweight force decreases to balance the governor spring force at the set position.

6. Reverse Thrust Reverse thrust can be obtained at any propeller speed. When the Power Lever is moved to the FULL REVERSE position it will increase compressor turbine speed (Ng) and propeller reverse pitch. The propeller governor is maintained in an under speeding condition in the reverse thrust range by controlling propeller speed with the Nf governing section of the propeller control.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT If Nf exceeds the desired speed the Nf governing orifice will open to decrease Py pressure in the computing section of the FCU and cause a reduction in fuel flow and Nf speed, thereby limiting the propeller speed and maintaining the CSU in an under-speed condition 7. Engine Shutdown The engine is shut down by moving the Fuel Condition Lever to the CUTOFF position. Fuel is returned to the pump inlet via the bypass passages. Fuel in the primary and secondary manifolds is drained via the dump valve ports in the flow divider and dump valve. Residual fuel is allowed to drain into the fuel drain can.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT ENGINE ACCESSORIES All engine-driven accessories with the exception of the propeller governor and propeller tachometergenerator are mounted on the accessory gearbox at the rear of the engine.

OIL PUMP The oil pump is located in the lowest part of the oil tank and is driven by the accessory gear shaft.

FUEL PUMP The engine driven fuel 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. Fuel from the oilfuel-heater enters the pump through a 74-micron filter. The pressure is boosted and enters the FCU through a 10-micron pump outlet filter. A bypass valve enables unfiltered fuel to reach the FCU in the event of a blockage. An internal passage originating at the FCU, returns bypass fuel from the FCU to the pump inlet downstream of the inlet screen.

Ng TACHO GENERATOR The Ng tacho-meter produces electric current which is used in conjunction with the gas generator RPM indicator.

TORQUE METER The torquemeter is located within the power reduction gearbox. It is a hydro-mechanical device connected to the first-stage reduction gear to provide an indication of engine output. It consists of a cylinder, piston and an oil metering-type plunger. Rotation of the first-stage reduction output drive ring gear is resisted by a helical spline which imparts an axial movement to the first stage ring gear and the torquemeter piston. This forces the piston onto the oil valve plunger allowing engine oil to enter the cylinder. This movement continues until the oil pressure within the torquemeter is equal to torque being applied to the first-stage ring gear.

A bleed hole is located in the torquemeter cylinder to allow a continuous flow of oil and to bleed pressure off when power is reduced. When engine oil is supplied to the plunger valve, it acts as a variable inlet metering orifice. The bleed hole acts a fixed calibrated leak. On acceleration, there is more oil supplied than bleeding away, so the pressure builds up in the torquemeter cylinder.

STARTER/GENERATOR The starter/generator is a 28-volt, 200 ampere engine driven unit. It functions as a motor for starting and, after engine start, as a generator for the electrical system. When operating as a starter, the speed sensing switch in the starter/generator will automatically shut down the starter, thereby providing overspeed protection and automatic shutoff. It is air-cooled by an integral fan and by ram air ducted from the front of the engine cowling.

ITT SENSING SYSTEM The ITT sensing system is designed to give the pilot an accurate indication of engine operating temperatures taken between the compressor and power turbines. It consists of 8 individual chromelalumel 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 of the power turbine vanes.

PROPELLER GOVERNOR See section 8.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT PROPELLER OVERSPEED GOVERNOR See section 8.

ENGINE FIRE DETECTION SYSTEM The engine fire detection system consists of a heat sensor in the engine compartment, a warning light on the annunciator panel and a warning horn.

The heat sensor consists of 3 flexible, closed loops. Each section of the loop is made up of a single wire surrounded by a continuous string of ceramic beads contained in an Inconel tube. The outer shell is connected to ground at the firewall and the wire inside is connected to the control box. The beads have a characteristic that causes them to lower their electrical resistance when the sensing element reaches its pre-set temperature value.

The core material in both elements prevents electrical current from flowing at normal temperatures.

When the elements are exposed to increasing temperatures, a current flow between the signal wire and ground. The control box detects a change in resistance in the loops and will illuminate the ENGINE FIRE light and trigger the warning horn.

The fire warning is initiated when temperatures in the engine compartment exceed 425° F (firewall), 625° F (around the exhaust) or 450° F (rear engine compartment). A test switch, labelled FIRE DETECT TEST, is located adjacent to the annunciator panel. When depressed, the ENGINE FIRE annunciator will illuminate and the fire warning horn will sound, indicating that the fire warning circuitry is operational. The system is protected by a “pull-off” type circuit breaker labelled FIRE DET, positioned third row and second from the left.

ENGINE REDUCTION GEAR SYSTEM The gearbox contains a two-stage planetary gear chain, three accessory drives and propeller shaft. The first-stage reduction gear is contained in the rear case, while the second stage reduction gear, accessory drives and propeller shaft are contained in the front case.

Torque from the power turbine is transmitted to the first-stage reduction gear, from there to the second-stage reduction gear and then to the propeller shaft. Reduction ratio is from a maximum power turbine speed of 33000 RPM down to a propeller speed of 1900 RPM. CHIP DETECTORS Two chip detectors are installed on the engine, one on the underside of the reduction gearbox and one on the underside of the accessory gearbox case. The chip detectors are electrically connected to an annunciator, labelled CHIP DETECTOR. The annunciator will illuminate when metal chips are present in one or both of the detectors. By itself this does not demand immediate action but if accompanied by signs of engine distress (erratic engine operation or fluctuation in engine power gage indications), engine operation may be continued at the discretion of the pilot consistent with crew safety.

OIL BREATHER DRAIN CAN Some aircraft have an oil breather can mounted on the right lower engine mount truss. This can collect any engine oil discharge coming from the accessory pads for the alternator drive pulley, starter/generator, air conditioner compressor and the propeller shaft seal. It should be drained after every flight.

The allowable quantity of oil discharge per hour of engine operation is 14 cc for aeroplanes with air conditioning and 11 cc for aeroplanes without. If the quantity of oil drained from the can is greater than specified, the source of the leakage should be identified and corrected.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN POWER PLANT

INTENTIONALLY LEFT BLANK

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN PROPELLER

PROPELLER The engine is equipped with a McCauley aluminium, constant speed, full-feathering, reversible, singleaction, governor controlled, three bladed propellers.

A setting is introduced into the governor with the propeller control lever. The propeller uses oil pressure which opposes the force of springs and counterweights to obtain the correct pitch for the engine load. Oil pressure from the propeller governor drives the blades toward low (fine) pitch (increasing RPM).

Springs and counterweights drive the blades toward high (coarse) pitch (decreasing RPM).

The source of oil pressure is furnished by the engine oil system, boosted in pressure by the governor gear pump and supplied to the propeller hub.

PROPELLER GOVERNOR The propeller governor is located in the 12 o'clock position on the front case of the reduction gearbox.

It combines the functions of a normal propeller governor (CSU), a reversing (BETA) valve and a power turbine (Nf) governor into a single unit.

CONSTANT SPEED MODE Under normal flight conditions, the governor acts as a CSU by maintaining propeller speed selected by the pilot. If the Power Lever is pushed forward, the FCU simply schedules an increase in fuel flow.

Thus, effectively more energy is available to turn the turbine. The turbine is forced to absorb the extra energy that is transmitted to the propeller in the form of torque. If there were no governor, the propeller speed would increase. Instead, the blade angle increases to maintain a constant propeller RPM and the propeller is allowed to take a larger bite of air, hence the increase in torque.

Therefore, the governor varies the propeller blade pitch angle to match the load to the engine torque in response to changing flight conditions. A spring-loaded pilot valve is installed in a driveshaft centre bore within the propeller governor. Ports in the driveshaft and the position of the pilot valve in the shaft, control the direction of oil flow within the housing. The rotating shaft with the rotating flyweights determines the position of the pilot valve while opposing spring load on the valve is varied by the speed adjusting lever at the head of the governor. The speed adjusting lever is connected to the Propeller Control Lever in the cockpit.

OPERATION A. On-Speed Cycle With the Propeller Control Lever set to a desired RPM and the propeller blade at the correct pitch angle to absorb the power developed by the engine, the centrifugal force of the rotating flyweights balances the force of the governor speeder spring. This condition positions the pilot valve so that the valve oil ports are closed and no oil flow occurs between the propeller oil pump and the propeller servo piston.

B. Under speed Cycle Under speeding occurs when the propeller RPM falls below the set RPM. Governor spring tension overcomes the reduced flyweight force and pivots them inwards, forcing the pilot valve downward and opening the oil ports.

This allows a flow of pressure oil from the governor pump to the propeller servo piston which, in turn, overcomes the combined forces of the propeller counterweights and return springs to reduce propeller blade pitch angle. This reduces the load and allows propeller RPM to increase.

This increased RPM is sensed by the governor flyweights which force the pilot valve upwards until

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN PROPELLER an equilibrium between speeder spring tension and centrifugal force of the flyweights is achieved causing the oil ports to close.

C. Overspeed Cycle In an overspeed condition, the governor flyweights pivot outward and overcome the governor speeder spring tension and raise the pilot valve. This uncovers the ports in the drive gear shaft and allows pressure oil in the propeller servo to flow into the reduction gearbox sump.

As the propeller blade angle increases, the load on the engine increases and propeller RPM decreases. This in turn causes the centrifugal force on the flyweights to decrease, allowing the speeder spring tension to return the flyweights to the vertical position, causing the pilot valve to cover the ports and block the oil supply from the servo piston.

LOW PITCH STOP During the approach to land when the Power Lever is closed and the Propeller Lever is in the maximum position the propeller blades are sitting on the low pitch stop at 15.6°. With the propeller blade angle and airspeed both reduced, the governor will tend to decrease the blade angle to restore the on-speed condition. If the blade angle could decrease all the way, the propeller would create so much drag that performance will be decreased. The low pitch stop is a mechanical device that prevents the blades from rotating from the lowest possible blade position to reverse. The low pitch stop is controlled by the power lever.

The propeller servo piston is connected by three spring loaded sliding rods to a feedback ring. A carbon brush block riding on the feedback ring transfers movement of the ring through the reversing lever to the BETA valve of the governor. As the propeller pitch angle continues to decrease, each flange on the propeller dome pushes the nut and the attached rod forward. As the rod moves forward, it pulls the feedback ring forward.

In turn the BETA valve moves forward and blocks off flow of oil to the propeller preventing any further pitch change. When the Power Lever is into the BETA range, the low pitch stop is automatically opened to allow the propeller blade angle to rotate up to the full reverse position.

BETA MODE The region between 18° and 5° blade angle is referred to as the BETA range. In this area of operation, the co-ordinated RPM/blade angle schedule that governs normal flight operations is no longer appropriate. As a result, whenever operating in the BETA range, the propeller blade angle is no longer being controlled by the propeller governor, but is directly controlled by the position of the Power Lever itself. The BETA range is there for a nongoverning range. In this range Ng remains at a value it had when the power lever was at IDLE. Beyond this range, the blade angle moves into reverse and the Ng increases.

REVERSE CONTROL The geometry of the power lever linkage through the cam box is such that power lever movement from idle to full forward thrust has no effect on the position of the BETA valve. However, when the power lever is moved into the reverse thrust range, the reversing lever pulls the BETA valve back.

Flow of oil to the propeller dome is re-established and the blades will move through low pitch and into a negative pitch range. As the propeller moves into low pitch, the feedback ring moves forward in an attempt to close the BETA valve and prevent any further pitch change. This cannot occur while the propeller reversing lever continues to move towards full reverse and pull the BETA valve back As the power lever is moved further into the reverse range, the blade angle decreases and the Ng simultaneously increases to approximately 85% - 88%. The power turbine governor (Nf) section of the propeller governor is simultaneously reset to approximately 95% RPM by the reset lever and

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN PROPELLER associated interconnecting rod. Selection of reverse thrust also repositions the FCU arm to increase fuel flow and consequently power output.

FEATHERING Moving the speed adjusting lever towards a pre-set feathering stop raises the pilot valve and decreases the oil pressure to the propeller servo piston. This decrease in oil pressure allows the piston to move under the influence of feathering and return springs, to rotate the propeller blades to, a positive coarse pitch or feathering position regardless of governor flyweight force acting on the pilot valve. The feathered blade angle is 88°.

UNFEATHERING Position the propeller lever forward of the feather gate. The unfeathering system uses oil pressure to force the propeller out of feather.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN PROPELLER POWER TURBINE (Nf) OVERSPEED GOVERNOR The function of the Nf governor section of the propeller governor, during normal forward thrust operation, is to protect the engine against a possible power turbine overspeed in the event of a propeller governor failure.

During reverse operation, the Nf governor is set below the propeller governor selected speed. This acts to control propeller speed via the FCU servo system and thus reduce the power generated by the gas generator to below required setting (approximately 5% less than the selected propeller speed).

In the event of a turbine overspeed, an air bleed orifice in the propeller governor is opened by flyweight action to bleed off compressor discharge pressure (Py) through the governor and computing section of the FCU. Compressor discharge pressure, acting on the FCU governor bellows, decreases and moves the metering valve in a closing direction.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN PROPELLER PROPELLER OVERSPEED GOVERNOR The propeller overspeed governor is installed in parallel with the propeller governor and is mounted at the 10 o’clock position on the front case of the reduction gearbox. It is incorporated to control any propeller overspeed condition by immediately bypassing pressure oil from the propeller servo to the reduction gearbox sump.

When an engine overspeed occurs, the increased centrifugal force sensed by the flyweights overcomes the spring tension, lifts the pilot valve and bypasses propeller pitch change mechanism oil back to the reduction gearbox. This permits the combined forces of the counterweights and return springs to move the blades toward a coarse pitch position, absorbing the engine power and preventing further engine overspeed.

A solenoid valve, which resets the governor to a value below its normal overspeed setting provides ground testing. The overspeed governor test switch is located on the left side of the instrument panel.

When depressed, a solenoid is actuated on the propeller overspeed governor. Speeder spring pressure is reduced and flyweights move out at a lower speed to simulate an overspeed condition. Propeller RPM is restricted when the power lever is advanced. During the test the propeller RPM should not exceed 1750 ± 60 RPM.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN PROPELLER

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN FUEL SYSTEM

FUEL SYSTEM COMPONENTS Two vented, integral fuel tanks with shutoff valves, a fuel selector 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, 14 fuel nozzle assemblies and a fuel drain can and drain.

SYSTEM Fuel flows from the tanks through two shutoff valves at each tank. 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.

A head of fuel is maintained around the ejector boost pump and auxiliary boost pumps which are contained within the reservoir. 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 is driven by motive fuel from the FCU. It normally provides fuel flow when the engine is operating. In the event of failure of the ejector boost pump, the electric auxiliary boost pump will automatically turn on, on condition that the boost pump switch is in the NORM position. The auxiliary boost pump is also used to supply fuel during starting.

Fuel in the reservoir manifold assembly flows through the fuel shutoff valves located on the aft side of the firewall. Fuel is then routed through the fuel filter located on the front side of the firewall. The filter incorporates a bypass feature.

Fuel is then routed through the oil-to-fuel heater to the engine driven pump 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 pressure and delivers it to the FCU via a 10-micron filter. A bypass valve and cored passages in the pump body enables unfiltered high-pressure fuel to flow to the FCU in the event the outlet filter becomes blocked.

The FCU meters the fuel and directs it to the flow divider. The flow divider schedules the metered fuel between the 10 primary and 4 secondary fuel manifolds. The result is an even, efficient spray pattern through all operational speeds. In the primary spin chamber, a change in direction puts a spinning motion on the fuel and establishes the proper spray angle and helps with atomisation. At that point the fuel is discharged from the primary nozzle. As the engine accelerates from start-up, the fuel pressure begins to increase. At approximately 90 psi, it causes the flow divider to open and part of the incoming fuel flow is channelled to the secondary spin chamber. From the nozzles the fuel is distributed to the combustion chamber.

When the fuel cut-off valve in the FCU closes during engine shutdown, both primary and secondary fuel nozzles are connected to a dump valve port. Residual fuel in the manifold drains into a fireproof can on the front left side of the firewall. This can should be drained daily.

Fuel venting is essential to system operation. A 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 from each tank. One vent from each tank protrudes from the trailing edge of the wing at the tips. The fuel reservoir is also vented to both wing tanks.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN FUEL SYSTEM FIREWALL SHUTOFF VALVE A manual shutoff valve is located on the aft side of the firewall. It enables the pilot to shut off all fuel from the reservoir to the engine. The shutoff valve is controlled by a red push-pull knob labelled FUEL SHUTOFF-PULL OFF is located on the pedestal. The push-pull knob has a press-to-release button in the center which locks the knob in position when the button is released.

FUEL TANK SELECTORS Two fuel tank selectors mechanically control the position of the fuel tank shutoff valves at each wing tank. When in the ON position, both shutoff valves in the tank are open, allowing fuel from that tank to flow to the reservoir. When the aircraft is parked on a slope, both fuel tank selectors should be in the OFF position. By leaving the fuel selector of the elevated wing in the OFF position only will not completely eliminate this cross flow. Over a prolonged period, there is enough seepage through the reservoir fuel vent to create a wing-heavy condition for the next flight.

AUXILIARY BOOST PUMP SWITCH The auxiliary boost pump switch is labelled FUEL BOOST and has OFF, NORM and ON position.

OFF POSITION The auxiliary boost pump is inoperative.

NORM POSITION The auxiliary boost pump is armed and will operate if the fuel pressure in the fuel manifold assembly drops below 4.75 psi. This switch position is used for all normal engine operations where main fuel flow is provided by the ejector boost pump and the auxiliary pump is used as a standby.

ON POSITION When the auxiliary boost pump switch is placed in the ON position, the pump operates continuously.

This position is used for the following : • • • •

Engine starting.

Any time the auxiliary boost pump cycles on and off with the switch in the NORM position.

Conditions of near fuel exhaustion.

For all operations using AVGAS.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN FUEL SYSTEM FUEL FLOW INDICATOR The fuel flow indicator indicates fuel consumption in pounds per hour based on Jet A fuel. Flow of fuel is measured downstream of the FCU just before being routed to the fuel divider. When battery power is removed, the indicator needle will stow below zero in the OFF position.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN FUEL SYSTEM FUEL QUANTITY INDICATORS Fuel quantity is measured by four transmitters in each tank and indicated by two electrically operated fuel quantity indicators. The indicators measure volume and are calibrated in pounds (based on the weight of Jet A fuel on a standard day 6,7 lbs/gal at 15°C) and gallons. An empty tank is indicated by a red line and the letter E. When showing empty, approximately 2,5 gallons remain in the tank as unusable fuel.

Because of the long fuel tanks, fuel quantity indicator accuracy is affected by uncoordinated flight or a sloping ramp if reading the indicators while on the ground. To obtain an accurate reading, ensure that the aircraft is parked laterally level, or if in flight, in a coordinated and stabilised condition.

ANNUNCIATORS FUEL LOW Two amber fuel low warning annunciators are labelled LEFT FUEL and RIGHT FUEL LOW. Each annunciator will illuminate when the fuel in the respective tank is less than 25 gallons.

RESERVOIR FUEL LOW A red reservoir fuel low annunciator labelled RESERVOIR FUEL LOW will illuminate when the level of fuel in the reservoir drops to approximately one-half full. There is only enough fuel in the reservoir for 1½ minutes of engine operation maximum continuous power after illumination.

FUEL PRESSURE LOW An amber fuel pressure low warning annunciator labelled FUEL PRESS LOW will illuminate when the fuel pressure in the reservoir fuel manifold assembly is below 4.75 psi.

AUX FUEL PUMP ON An amber auxiliary fuel pump on annunciator labelled AUX FUEL PUMP ON will illuminate when the auxiliary boost pump switch is in the ON position or when the boost pump switch is in the NORM position and the fuel pressure in the fuel manifold assembly drops below 4.75 psi.

FUEL DRAIN VALVES

Fuel drain valves are located at the lower surfaces of each wing at the inboard end of the tank.

Outboard fuel tank drain valves may be installed and their use is recommended if the aircraft is parked with one wing low on a sloping ramp. The drain valves are constructed so that a Phillips screwdriver can be used to depress the valve and then twist to lock it in the open position.

The drain valve for the reservoir consists of a recessed T-handle. When pulled out, fuel from the reservoir drains out the rear fuel drain pipe located adjacent to the drain valve. The drain valve for the fuel filter consists of a drain pipe which can be depressed upwards to drain fuel from the filter. If contamination is detected, drain all fuel drain points again. Take repeated samples of all fuel drain points until all contamination has been removed.

FUEL DRAIN CAN When the engine is shut down, residual fuel in the engine drains into a fuel drain can mounted on the front left side of the firewall. The can should be drained once a day or at intervals not exceeding 6 shutdowns. A drain valve on the bottom side of the cowling is provided to drain the fuel into a suitable container.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN FUEL SYSTEM FUEL PUMP DRAIN RESERVOIR To control expended lubrication oil from the engine fuel pump drive coupling area and provide a way to determine if fuel is leaking past the fuel pump seal a drainable reservoir collects allowable discharge of oil and any fuel seepage. The reservoir is mounted on the front left side of the firewall. It should be drained once a day or at intervals not exceeding 6 shutdowns. A quantity of up to 3 cc of oil and 20 cc of fuel discharge per hour of engine operation is allowable. In addition, a FCU bearing failure will be indicated by a blue dye in the expended oil. This requires immediate attention and under no circumstances should the aircraft be flown before this situation is corrected.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN ELECTRICAL SYSTEM

ELECTRICAL SYSTEM The aircraft is equipped with a 28-volt dc electrical system. The system uses a 24-volt battery as a source of electrical energy and a dual-purpose starter-generator. An optional standby alternator is available for the use as a standby source in the event of the main generator system malfunctioning.

ELECTRICAL BUSES Power is supplied to the electrical and avionics circuits through 2 general, 2 avionics and a battery bus. The battery bus is energised continuously for the memory keep alive, clock and cabin/courtesy lights. The two general buses are on any time the battery switch is turned on. All DC buses are on any time the battery switch and two avionics switches are turned on.

STARTER-GENERATOR The starter-generator is mounted on the engine accessory gearbox and is driven by the engine through a splined shaft. It is cooled by an integral fan and a blast tube located above the right forward cowling.

It functions as a motor during engine start and as a generator after starting.

In the starter mode it is powered by either the battery or an external power source. After engine start, the generator output is 28 volt and 200 amp. When operating as a generator, it supplies power to operate the electrical systems and maintains the battery's state of charge. The unit is controlled by a generator control unit.

GENERATOR CONTROL UNIT (GCU) The GCU is mounted inside the cabin on the left forward fuselage sidewall. It provides the electrical control functions necessary for the operation of the starter-generator. It provides automatic starter cut-out at 46% Ng. Below 46% the starter-generator functions as a starter. Above 46% it functions as a generator when the starter switch is OFF. The GCU provides voltage regulation plus high voltage protection reverse current protection. A rheostat increases the resistance in the field circuit and less current flows through the field winding resulting in a decrease in the strength of the magnetic field. The rotating armature now turns in a weaker field with the result of a lower generator output voltage.

A differential relay switch protects the generator from reverse current from battery voltage. It is essentially an on/off switch that is controlled by the difference in voltage between the battery bus and the generator output. The differential relay switch connects the generator to the electrical system’s main bus whenever generator voltage exceeds bus voltage by at least 0.5 volt. If on the other hand, bus voltage exceeds generator output voltage, the reverse current relay opens and takes the generator offline. In the event of the above, the generator is automatically disconnected from the buses and the red GENERATOR OFF annunciator illuminates.

GROUND POWER MONITOR The ground power is located inside the electrical power assembly mounted on the left side of the firewall. This unit senses voltage applied to the external power receptacle and closes the external power contactor when the applied voltage is within limits (24 - 28 volts and 800 - 1700 amp).

It also senses airplane bus voltage and will illuminate the VOLTAGE LOW annunciator when the bus voltage drops to 24.5 volts.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN ELECTRICAL SYSTEM GROUND SERVICE PLUG RECEPTACLE The external power receptacle is installed on the left side of the engine compartment near the firewall.

It permits the use of an external power source for starting and lengthy maintenance on the left side of the engine compartment.

External power control circuitry is provided to prevent external power and the battery from being connected together during starting. The ground service circuit incorporates polarity reversal and over-voltage protection.

EXTERNAL POWER SWITCH The external power is a guarded toggle-type switch with the following positions: OFF POSITION Battery power is provided to the main bus and to the starter circuit. External power cannot be applied to the main bus. With the generator switch in the ON position, power is applied to the generator control circuit.

STARTER POSITION External power is applied to the starter circuit only. Battery power is applied to the main bus. No generator power is available in this position. If external power drops off during the start cycle, the external power switch must be placed in the OFF position to reconnect the battery to the starter if motoring is required.

BUS POSITION External power is applied to the main bus. No power is available to the starter. The battery, if desired, can be connected to the main bus and the external power by the battery switch. Battery charge should however be monitored to avoid overcharge.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN ELECTRICAL SYSTEM BATTERY The system uses a 24-volt lead-acid battery with a capacity of 45 amp-hr or a 24 volt Ni-cad battery with a 40 amp-hr capacity. The Ni-cad battery has a longer service life, high discharge and short recharge capability.

THERMAL RUNAWAY If for any reasons a cell’s temperature increases, its voltage and internal resistance decreases. Three factors lead to increasing a cell’s temperature : • • •

Excessively high discharge rate.

Excessively high ambient conditions, particularly with a poorly or improperly ventilated battery compartment.

Deterioration of the plates’ separator material, which allows oxygen from the positive plate to interact with a negative plate where it will chemically interact with the cadmium and generate heat.

The beginning of the problem is an excessive discharge rate. On short legs in IMC at night, the battery is constantly being drained and recharged, which generates excessive battery temperature. Due to the way individual cells are installed in the battery case, the outer cells dissipate heat through the sides of the case and as a result run slightly cooler than the middle cells. As the battery temperature rises, it is the inner cells that get the hottest.

The generator attempts to recharge the battery by supplying sufficient current to the cells. As the cell temperature increases, its resistance decreases simply helping the generator. The excessive current causes the cells to overheat even more, further reducing their resistance. The porous plastic strip between the individual plates breaks down, further complicating the problem. At this point, shutting off the generator will stop the problem.

Meanwhile, the inner cell’s internal resistance and voltage becomes so low that the voltage of the good cells surrounding it will be higher by comparison and they will begin to feed the bad cells. When any cell begins to receive voltage from a surrounding cell, isolating the battery from the generator will no longer have any effect; the battery is self-destructing. The only remaining solution is to land as soon as possible.

BATTERY HOT light This amber light illuminates when the battery temperature is 140 - 160°C. BATTERY OVERHEAT light This red light illuminates when the temperature exceeds 160°C. Immediately turn the battery switch OFF.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN ELECTRICAL SYSTEM BATTERY SWITCH In the ON position battery power is supplied to the two general buses. The OFF position cuts power to all buses except the battery bus.

GENERATOR SWITCH The generator has ON, RESET and TRIP positions.

RESET AND TRIP These positions are momentary and are spring loaded back to ON position. If the GENERATOR OFF or VOLTAGE LOW lights illuminate, place the switch momentarily in the RESET position to restore generator power. If erratic operation is observed, the system can be shut off by placing the switch momentarily in the TRIP position. After a suitable waiting period, generator operation may be recycled by placing the generator switch momentarily to RESET. ON POSITION The GCU will automatically control the generator line contactor for normal generator operation.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN ELECTRICAL SYSTEM AVIONICS POWER SWITCHES The avionics power switches control power to number 1 or number 2 avionics buses. They should be in the OFF position prior to turning the battery switch ON or OFF, engine starting or applying external power. All avionics may be turned on or off by operating the AVIONICS power switches rather than by operating all of the individual avionics equipment switches.

AVIONICS BUS TIE SWITCH The avionics bus tie switch is guarded in the OFF position. It connects the number 1 and number 2 avionics buses together in the event of failure of either bus feeder circuit. Since each avionics bus is supplied power from a separate current limiter on the power distribution bus, failure of the current limiter can cause failure of the affected bus. Placing the bus tie switch to the ON position will restore power to the failed bus. Operation without both bus feeder circuits may require an avionics load reduction.

CIRCUIT BREAKERS Most of the electrical circuits are protected by “pull-off” type circuit breakers. The circuit breaker panel is mounted on the left sidewall. Once a circuit breaker has popped, allow it to cool for approximately 3 minutes before resetting. A circuit breaker should be reset once only and at the discretion of the pilot consistent with crew safety.

Ensure that all circuit breakers are in before all flights. Never operate with disengaged circuit breakers without a thorough knowledge of the consequences.

VOLT/AMMETER A volt/ammeter and a four-position rotary-type selector switch is mounted on the left side of the instrument panel so that the electrical system can be monitored. The selector switch has GEN, ALT, BATT and VOLT positions.

GEN POSITION Measures generator current. The meter is connected to the generator shunt and will display current in amps flowing from the starter/generator to the distribution bus.

ALT POSITION Current from the standby alternator to the number 1 and 2 bus power circuit breakers is displayed.

BATT POSITION Battery charge rate or discharge is measured because the meter displays both positive and negative currents.

VOLT POSITION System voltage is measured from either the battery or the generator.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN ELECTRICAL SYSTEM STANDBY ELECTRICAL SYSTEM A standby electrical system is installed in the event the main generator system malfunctions in flight.

This system includes an alternator at a 75-amp capacity rating. The alternator is belt driven from the accessory pad on the rear of the engine.

Initially battery power is required to excite the magnetic field. Field excitation to the alternator control is supplied through diode logic from either a circuit breaker in the standby alternator assembly or KEEP ALIVE 2 circuit breaker in the main power relay box. A diode allows electron flow in one direction only. After alternator operation is initiated, the alternator is self-excited.

System monitoring is provided by two amber STBY ELEC PWR ON and STBY ELEC PWR INOP annunciator lights. Supplied amperage can be monitored on the voltammeter by placing the switch to ALT position.

The STBY PWR switch is turned on after start. The standby power system is now armed and will automatically supply power to the main bus only if the system voltage drops below 27.5 volts.

In order to utilise the 75-ampere capacity from the standby alternator, the AVIONICS STBY PWR and AVIONICS BUS TIE switch/breaker must be turned ON. AVIONICS 1 and 2 switches should be turned OFF to avoid connecting the standby power system to a possible fault in the primary power when operating on standby power. Power to the two main buses is limited to 30 amps per bus. The load may have to be reduced to prevent battery discharge. In the event of a fault in the primary relay box, the primary power supply system can be isolated by pulling the six 30-amp bus feeder circuit breakers.

Immediately after start, the STBY ELECT PWR ON annunciator may illuminate, indicating the standby electrical system is operating to help replenish the electrical power used for starting. The

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN ELECTRICAL SYSTEM STBY ELECT PWR INOP annunciator will not illuminate except in the event of a broken alternator drive belt or an electrical malfunction in the standby electrical system.

The standby alternator receives field current from the KEEP ALIVE 2 circuit. In an emergency the standby alternator can be brought on line without turning the battery switch on. Normal engine shutdown procedures call for the standby power to be switched off prior to shutting the engine down and turning the battery switch off. If the standby power switch is inadvertently left on, several red lights in the annunciator panel will remain illuminated after the battery is turned off.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN LIGHTING

LIGHTING EXTERIOR LIGHTING All exterior lights are controlled by toggle switches located on the lighting control panel on the left side of the instrument panel. The switches are ON in the up position and OFF in the down position.

NAVIGATION LIGHTS Conventional navigation lights are installed on the wingtip and the tail cone stinger. The lights are protected by a circuit breaker labelled NAV LIGHT. LANDING LIGHTS Two landing lights are mounted in the outboard section of each wing leading edge. They provide illumination forward and downward during take-off and landing. The lights are protected by two circuit breakers labelled LEFT LDG LT and RIGHT LDG LT. The landing lights have a relatively short service life and should be used for take-off and landing only.

TAXI/RECOGNITION LIGHTS Two taxi/recognition lights are mounted inboard of each landing light. These lights are focused to provide illumination of the area forward of the aircraft during ground operation. The lights are also used in the traffic pattern or enroute. They are protected by a circuit breaker labelled TAXI LIGHT. STROBE LIGHTS A strobe lights with remote power supply is installed on each wingtip. The lights are used to enhance anti-collision protection and should be turned OFF when taxiing. Do not operate the strobe lights in conditions of fog, cloud or haze as the light beam can cause vertigo. The lights are protected by a circuit breaker labelled STROBE LIGHT. FLASHING BEACON LIGHT A red flashing light is installed on top of the vertical fin. The light is visible through 360°. The light is protected by a circuit breaker labelled BEACON LIGHT. Do not operate the flashing beacon when flying through clouds as the reflected light can cause vertigo.

COURTESY LIGHTS A courtesy light is installed under each wing. The lights illuminate an area adjacent to the crew entry doors. The lights operate in conjunction with the cabin lights and are controlled by the cabin light switches.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN LIGHTING INTERIOR LIGHTING CONTROL KNOBS Instrument and control panel lighting is provided by integral, flood and post lights. Four concentrictype dual lighting control knobs are grouped together on the lower part of the instrument panel.

Controls vary the intensity of the instrument panel lighting, the left sidewall switch and circuit breaker lighting, the pedestal lighting and the overhead panel lighting.

MAP LIGHT A map light is mounted on the bottom of the pilot's control wheel. Brightness is adjusted by means of a rheostat control knob on the bottom of the control wheel.

CABIN LIGHTS Four cabin lights are installed in the interior of the airplane. These lights are controlled either by a two-position toggle switch, labelled CABIN, on the lighting control panel or rocker-type switches located on the inside sidewall panel just forward of the cargo door and the passenger entry door. This light circuit does not require power to be applied to the main electrical system buses for operation.

The courtesy lights circuit may be equipped with a solid-state timer which allows the lights to remain illuminated for period of 30 minutes.

PASSENGER READING LIGHTS Passenger reading lights may be installed near each of the aft passenger positions. The lights are located in small convenience panels above the seats. A pushbutton-type ON, OFF switch, mounted in each panel, controls the lights.

NO SMOKING/FASTEN BELT SIGN This installation consists of a small lighted panel mounted in the cabin headliner above the right side of the forward cabin area. The lights are controlled by two toggle-type switches, labelled SEAT BELT and NO SMOKE. These warning sign lights are protected by a circuit breaker labelled SEAT BELT SIGN.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AIR CONDITIONING

AIR CONDITIONING VENTILATING AIR Ventilating air is obtained from an inlet on each side at the forward fuselage and through two ram air inlets, one on each wing at the upper end of the struts. The wing inlet ventilating air is routed through the wing into the plenum chamber located in the centre of the cabin top. The plenum chamber distributes the air to the individual overhead outlets at each seat position.

VENTILATION FAN SYSTEM The ventilation fan system provides supplemental cabin ventilation. The system controls are located on the overhead console and consists of two rotary-type vent air controls, labelled VENT AIR. The vent air controls operate shutoff valves in the left and right wing to control the flow of ram ventilating air which enters the inlet in each upper wing strut fairing. The vent air controls also operate switches in two ventilation fan circuits to control fan operation. When the vent air controls are rotated to the CLOSE position, the shutoff valves are closed. Rotating the knobs beyond the OPEN to FAN position, a mechanism on the control actuates a switch to operate the duct-mounted fan just inboard of the shutoff valve in each wing. The switch will not operate the fan until the shutoff valve is open, thus assuring a supply of cool air to the fan motor. Whenever the vent air controls are in the OPEN position, ram airflow is ducted to the overhead ventilating outlets. This airflow can be augmented by rotating the controls to the FAN position.

System electrical protection is provided by two “pull-off” type circuit breakers labelled LEFT VENT BLWR and RIGHT VENT BLWR. INSTRUMENT PANEL VENT KNOBS Two vent knobs, labelled VENT, PULL ON, are located on each side of the instrument panel. Each knob controls the flow of ventilating air from an outlet located adjacent to each knob. Pulling each knob opens a small air door on the fuselage exterior.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AIR CONDITIONING CABIN HEATING SYSTEM Hot compressor outlet air is routed from the engine through a flow control valve. This hot air is then mixed with cabin return air or warm compressor bleed air in the mixer/muffler. Air is then routed to the cabin distribution system. Controls direct the heated air to forward and/or aft portions of the cabin and to the windshield for defrosting.

BLEED AIR HEAT SWITCH The BLEED AIR HEAT switch is located on the cabin heat switch and control panel. The switch controls the operation of the bleed air flow control valve. The ON position of the switch opens the flow control valve, allowing hot bleed air to flow to the cabin heating system. The OFF position closes the valve, shutting off flow of air to the bleed air heating system.

TEMPERATURE SELECTOR KNOB A rotary temperature selector knob, labelled TEMP, is located on the cabin heat switch and control panel. This selector modulates the opening and closing action of the flow control valve to control the amount and temperature of air flowing into the cabin. Clockwise rotation of the knob increases the mass flow and temperature of the air.

If more cabin heat in needed on the ground, move the fuel condition lever to HIGH IDLE and/or select the GRD position of the mixing air control. For best results turn the selector to the full clockwise position and then slowly counter-clockwise to decrease the bleed airflow to the desired amount.

A temperature sensor, located in the outlet duct from the mixer/muffler operates in conjunction with the temperature selector knob. In the event of an overheat in the outlet duct, a temperature sensor will be energised, closing the flow control valve and shutting off the source of hot bleed air from the engine.

MIXING AIR PUSH-PULL CONTROL A push-pull control, labelled MIXING AIR, GRD-PULL, FLT-PUSH, is located on the cabin heat switch and control panel.

A. GRD Position With the push-pull control in the GRD position (pulled out), warm compressor bleed valve air (P2.5) is mixed with hot compressor outlet air (P3) in the mixer/muffler. This mode is used for ground operation when warm compressor bleed air (P2.5) is available. P2.5 is only available below 92% Ng. It can be used to augment hot compressor outlet bleed air (P3) during periods of cold ambient temperature.

B. FLT Position With the push-pull control in the FLT position (pushed in) cabin return air is mixed with hot compressor outlet air (P3) in the mixer/muffler. This recirculation of cabin return air enables the heating system to maintain the desired temperature.

If desired, the FLT position can be used on the ground when ambient temperatures are mild and maximum heating is not required. In this mode, excess warm compressor bleed valve air (P2.5) available at power settings below 92% Ng, is exhausted overboard from the mixing air valve.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AIR CONDITIONING

The mixing air push-pull control valve should always be in the FLT position in flight. Cabin return air must be allowed to flow through the mixing valve and blend with hot compressor outlet air during high engine power operation otherwise the system may overheat and cause automatic shutdown.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AIR CONDITIONING AFT /FORWARD CABIN PUSH-PULL CONTROL A push-pull control, labelled AFT CABIN-PULL, FWD CABIN-PUSH, is located on the cabin heat switch and control panel.

With the control in the AFT (pulled out) position, heated air is directed to aft cabin heater outlets located in the floor directly behind the pilots in the Cargo Version and on the cabin sidewalls at floor level for the Passenger Version.

With the control in the FWD CABIN position (pushed in), heated air is directed to the forward cabin through four heater outlets located behind the instrument panel and/or two windshield defroster outlets. The controls can be positioned at any intermediate setting desired for proper distribution of heated air to the forward and aft cabin areas.

DEFROST/FORWARD CABIN PUSH-PULL CONTROL A push-pull control, labelled DEFROST-PULL, FWD CABIN-PUSH, is located on the cabin heat switch and control panel.

With the control in the DEFROST position (pulled out), forward cabin air is directed to two defroster outlets located at the base of the windshield. The aft/fwd push-pull control must also be pushed in for availability of forward cabin air for defrosting.

With the control in the FWD CABIN position (pushed in), heated air will be directed to the four heater outlets behind the instrument panel.

CABIN HEAT FIREWALL SHUTOFF KNOB A push-pull shutoff knob, labelled CABIN HEAT FIREWALL SHUTOFF, PULL OFF, is located on the lower right side of the pedestal. When pulled out, the knob actuates two firewall shutoff knobs to the off position. One in the bleed-air supply line to the cabin heating system and the other in the cabin return air-line. This knob should normally be pushed in unless a fire is suspected in the engine compartment.

Do not place this knob in the OFF position when the mixing air control is in the GRD position. This will result in a compressor stall at low power settings when the compressor bleed valve is open. The engine must be shut down to relieve back pressure on the valves prior to opening the valves.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN OXYGEN

OXYGEN Oxygen is supplied by a 116.95 cubic foot capacity oxygen cylinder located in the fuselage tail cone.

The oxygen is under a pressure of 1850 psi. Below 200 psi flow rates are not predictable.

Cylinder pressure is reduced to 70 psi by pressure regulator attached to the cylinder. A shutoff valve forms part of the regulator assembly. An altitude compensating regulator, located between the pressure regulator and the oxygen supply lines, varies the flow of oxygen.

Partial re-breathing type oxygen masks are equipped with vinyl plastic hoses and flow indicators.

Hoses are high-flow type and colour coded with a blue band adjacent to the plug-in fitting.

Shutoff valves are located in the overhead consul. The valve shuts off the supply when supply not in use. It is mechanically connected to the shutoff valve on the cylinder.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN INSTRUMENTS

INSTRUMENTS PITOT-STATIC SYSTEM The pitot-static system supplies ram air pressure to the airspeed indicator and static pressure to airspeed indicator, vertical speed indicator and altimeter. The system is composed of a heated pitotstatic tube mounted on the leading edge of the left wing, a static pressure alternate source valve located below the de-ice/anti-ice switch panel, a drain valve located on the left sidewall beneath the instrument panel, an airspeed pressure switch located behind the instrument panel and associated plumbing.

The pitot-static heat system consists of a heating element in the pitot-static tube, a two-position toggle switch, labelled PITOT/STATIC HEAT and a “pull-off” type circuit breaker. When the switch is turned on, the element in the pitot-static tube is heated electrically to maintain proper operation in possible icing conditions.

A static alternate source is installed below the de-ice/anti-ice switch panel and can be used if the static source is malfunctioning. This valve supplies static pressure from inside the cabin instead of from the pitot-static tube. If erroneous instrument readings are suspected due to water or ice in the pressure line to the static pressure source, the alternate source valve should be pulled on. The pressure inside the cabin will vary with vents open or closed. Cabin static is usually lower; causing instruments to overread.

A drain valve is located on the cabin sidewall below the instrument panel. The valve is used to drain suspected moisture in the system by lifting the drain valve lever to the OPEN position. The valve must be returned to the CLOSED position prior to flight.

An airspeed pressure switch in the pitot- static system is used to actuate an airspeed warning horn.

The horn is located behind the headliner in the area above the pilot. It sounds when the airspeed exceeds 175 KIAS. A warning will also be heard in the pilot’s headset.

A second, independent pitot-static system is included for the right flight instrument panel.

VACUUM SYSTEM A vacuum system provides the suction necessary to operate the left hand attitude indicator and the right hand directional indicator. Vacuum is obtained by passing regulated compressor outlet bleed air through a vacuum ejector. Bleed air flowing through an orifice in the ejector creates the suction to operate the instruments. The vacuum system consists of the bleed air pressure regulator, a vacuum ejector on the forward left side of the firewall, a vacuum relief valve and vacuum system air filter, vacuum operated instruments, a suction gage and a vacuum-low warning annunciator.

SUCTION GAGE The suction gage is calibrated in inches of mercury. The desired suction ranges are 4.5 to 5.5 up to 15 000 feet, 4.0 to 5.5 from 15 000 to 20 000 feet, 3.5 to 5.5 from 20 000 to 25 000 feet. The 15K, 20K, 25K and 30K markings at the appropriate step locations indicate the altitude at which the lower limit of that arc segment is acceptable.

A suction reading out of these ranges may indicate a system malfunction or improper adjustment. In this case, the attitude and directional indicators should not be considered reliable.

VACUUM-LOW WARNING ANNUNCIATOR A red VACUUM LOW warning annunciator will illuminate when the suction is below 3.0 in. Hg.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN INSTRUMENTS ELECTRICALLY-OPERATED GYROS The right-hand horizontal situation indicator and the left-hand directional indicator have electrically operated gyros. These instruments are energised anytime the battery switch is ON and the circuit breakers are in. If take-off is soon after engine start, cage the gyro immediately after start by pulling the knob for approximately 5 seconds before releasing it. If time between start and take-off is 10 minutes or more, allow the gyro to run as caging is not necessary.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN ICED AND RAIN PROTECTION

ICE AND RAIN PROTECTION The flight into known icing equipment includes the following : • • • • • •

Pneumatic de-icing boots on the wings, wing struts, horizontal and vertical stabiliser leading edges.

Electrically heated propeller blade anti-ice boots.

A detachable electric windshield anti-ice panel.

A pitot-static heat system and heated stall warning system.

An ice detector light.

Engine inertial separator.

PNEUMATIC DE-ICE BOOTS The system components include a pressure line which leads from the engine bleed-air system pressure regulator to the vacuum ejector, three flow control valves and pressure switches, a timer, a system switch and circuit breaker, an annunciator and supply lines and pneumatically operated surface deice boots.

Boots expand and contract using pressure from the engine bleed-air system to the flow control valves.

Normally the flow control valves are open to create a vacuum to hold the boots against the leadingedge surfaces. When a de-icing cycle is initiated, the control valves are closed in order to remove the vacuum and allow the pressure to inflate the boots.

When the switch is held in the AUTO (upper) position and released one de-ice cycle will be activated.

The switch is off when placed in the middle position. Boots inflate according to the following sequence : • • •

Horizontal and vertical stabilizer boots.

Inboard wing.

Outboard wing and wing strut boots.

The total time required is 18 seconds and each sequence lasts 6 seconds. In the event of a malfunction of the timer, causing erratic operation of a sequence of a cycle, the switch can be held momentary in MANUAL (lower) position. This results in simultaneous inflation of all the boots. The system can be stopped (deflating the boots) by pulling the circuit breaker labelled DE-ICE BOOT. The pressure indicator annunciator illuminates within 3 seconds after initiating a cycle and remains on 3 additional seconds to the end of the first sequence. Through each of the remaining two sequences of the cycle, the annunciator will remain off during pressure build-up for about 3 seconds and then illuminate for about 3 seconds. When illuminated the pressure in the boot is about 15 psi. The absence of illumination during any one of the three sequences of a cycle indicates insufficient pressure for proper boot inflation and effective de-icing ability.

In rime ice conditions wait until approximately 1/2 to 3/4 inch of ice has accumulated. In clear ice conditions wait until approximately 1/4 to 3/8 inch of ice has accumulated. This procedure is recommended due to the high drag penalties associated with clear ice shapes. Do not cycle the boots during the approach and landing since inflation may increase the stall speeds by much as 10 kts. Do not use more than 20° flaps with heavy ice accumulations on the horizontal stabilizer leading edge.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN ICED AND RAIN PROTECTION PROPELLER ANTI-ICE BOOTS The system is operated by a three-position toggle switch labelled PROP. In the AUTO position electric current flows to an anti-ice timer which cycles the current simultaneously to the heating elements in the anti-ice boots on the three propeller blades in intervals of 90 seconds on and 90 seconds off.

The MANUAL position can be used in the event of failure of the anti-ice timer. The switch can be held in the MANUAL position to achieve emergency propeller anti-icing. Hold the switch in the lower position for 90 seconds.

The operation can be monitored on the ammeter labelled PROP ANTI-ICE AMPS. The ammeter should indicate 20 to 24 amps. An oil-operated pressure switch installed in the electrical circuit prevents the anti-ice system from being turned on without the engine running. A failure of this switch will be undetected unless the ammeter is monitored continuously. The PROP ANTI-ICE CONT circuit breaker protects the control circuit. The PROP ANTI-ICE circuit breaker protects the heater circuit.

WINDSHIELD ANTI-ICE PANEL A detachable, electrically heated windshield anti-ice panel can be mounted to the base of the pilot's windshield utilising a spring-loaded quick-release pin. Windshield anti-icing is controlled by a threeposition toggle switch labelled W/S. In the AUTO position electric current regulated by a controller flows to the anti-ice panel. In the event of malfunction in the system controller circuitry, the switch can be held in the MANUAL position to achieve windshield anti-icing.

ICE DETECTOR LIGHT An ice detector light is flush mounted near the upper left corner of the windshield. The switch is springloaded to the off position and must be held in the ON position to illuminate the light.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AUTOFLIGHT SYSTEM AND AVIONICS

AUTOFLIGHT SYSTEM AND AVIONICS THE KFC-150 AUTOPILOT SYSTEM The 150 AFCS is certified as a 2 axis autopilot control system. A third axis autopilot control for yaw is available as an option. The system incorporates an electric pitch trim system which provides auto trim during autopilot operation and manual electric trim for the pilot. Trim faults are visually and aurally enunciated. A lockout device prevents autopilot engagement until the system has been successfully pre-flight tested.

The following conditions will cause the autopilot to automatically disengage : • • • • • •

Electrical power failure.

Internal flight control system failure.

A loss of a valid compass signal (HDG flag) when a mode using heading information is engaged.

With the HDG flag present the autopilot may be reengaged in the basic wings-level mode with any vertical mode.

Roll rates in excess of 14° per second except when the CWS switch is held depressed.

Pitch rates in excess of 6° per second except when the CWS button is held depressed.

Actuating the manual electric trim.

DEFINITIONS AND TERMS A. Autopilot This autopilot system incorporates electrically-driven actuators that fly the aircraft by moving the control surfaces which are the ailerons, rudder (when optional yaw damper is installed), elevators and the elevator trim.

B. Flight Director The flight director system incorporates a panel-mounted computer that calculates intercept angles and displays them to the pilot as pitch and steering recommendations on the Flight Command Indicator (FCI). The recommended pitch and steering display serve as a reminder to the pilot as to which way he should fly the aircraft to get the desired results of his mode selector input.

The flight director may be used with the autopilot engaged or disengaged. In the latter case, the pilot manually flies the aircraft to satisfy the command bar on the FCI which is positioned by the computer rather than allowing the autopilot to satisfy the computed commands.

• • •

The autopilot can only be coupled to the NAV 1 receiver The flight director (FD) mode must be selected before the autopilot engage mode (AP ENG) can be selected.

The AVIONICS POWER #1 switch supplies power to the autopilot and ELEV TRIM circuit breakers.

THE AUTOPILOT AND FLIGHT DIRECTOR COMPUTER A. Yaw Damper (Yd) Illuminates when the yaw damper (optional) is engaged.

B. Mode Annunciators (Fd, Alt, Hdg, Nav, Apr and Bc) Illuminate when a mode is selected by the corresponding mode selector button.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AUTOFLIGHT SYSTEM AND AVIONICS C. Glide Slope (GS) Annunciator Illuminates continuously when the autopilot is coupled to the glide slope signal. The GC annunciator will flash if the glide slope signal is lost (GS flag in CDI). The autopilot reverts to pitch hold operation. If a valid glide slope signal returns within 6 seconds, the autopilot will automatically recouple in the GS mode. If the valid signal does not return within 6 seconds, the autopilot will remain in pitch attitude hold until such time that a valid glide slope returns and the aircraft passes through the glide slope. At that point, GS couple will re-occur.

D. Trim Warning Light (Trim) Illuminates continuously whenever trim power is not on or the system has not been pre-flight tested. The TRIM warning light illuminates and an audible warning will sound whenever a trim fault is detected. The auto trim system is monitored for the following failures : • • •

Trim servo running without a command.

Trim servo not running when commanded to.

Trim servo running in the wrong direction.

E. Autopilot (Ap) Annunciator Illuminates continuously when the autopilot is engaged. Flashes 12 times whenever the autopilot is disengaged (an aural alert will sound for 2 seconds).

F. Autopilot Engage (Ap Eng) Button When pushed, engages the autopilot when all logic conditions are met.

G. Pre-flight Test (Test) Button When momentarily pushed, initiates pre-flight test sequence which : • • • • • • •

Automatically turns on all lights.

Tests the roll and pitch rate monitors.

Tests the auto trim fault monitor.

Checks the manual trim drive voltage.

Tests all autopilot valid and dump logic.

If the pre-flight is successfully passed, the AP annunciator light will flash for 6 seconds (an aural tone will also sound simultaneously with the annunciator flashes).

The autopilot cannot be engaged until the autopilot pre-flight test are successfully passed.

H. Back Course Approach (Bc) Mode Selector Button When pushed, selects Back Course Approach mode. This mode functions identically to the approach mode except that response to the LOC signals is reversed. Glide slope coupling is inhibited.

I.

Approach (Apr) Mode Selector Button When pushed, selects Approach mode. This mode provides all angle intercept, automatic beam capture and tracking of VOR, RNAV or LOC signals plus glide slope coupling in case of an ILS. The tracking gain of the APR is greater than the gain in NAV mode. The APR annunciator will flash until the automatic capture sequence is initiated. On the Remote Mode Annunciator, APR ARM will annunciate until the automatic capture sequence is initiated. At beam capture, APR CPLD will annunciate.

J.

Navigation (Nav) Mode Selector Button When pushed, selects NAV mode. This mode provides all angle intercept (with HSI), automatic beam capture and tracking of VOR, RNAV or LOC. The NAV annunciator will flash until the automatic capture sequence is initiated. On the Remote Mode Annunciator, NAV ARM will

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AUTOFLIGHT SYSTEM AND AVIONICS annunciate until the automatic capture sequence is initiated. At beam capture, NAV CPLD will annunciate.

K. Heading (Hdg) Mode Selector Button When pushed, selects the heading mode which commands the aircraft to turn to and maintain the heading selected by the heading bug on the HSI. A new heading may be selected at any time and will result in the aircraft turning to the new heading with a maximum bank angle of about 20°.

Selecting HDG mode will cancel NAV, APR or BC track modes.

L. Altitude Hold (Alt) Mode Selector Button When pushed, selects altitude hold mode which commands the aircraft to maintain the pressure altitude existing at the moment of selection. Engagement may be accomplished in the climb, cruise or descent. In the APR mode, altitude hold will automatically disengage when the glide slope is captured.

M. Flight Director (Fd) Mode Selector Button When pushed, selects the flight director mode, bringing the command bar in view on the AI and will command wings-level and pitch attitude hold. The FD must be selected prior to autopilot engagement.

N. Vertical Trim Control A spring-loaded-to-center rocker switch will provide up or down pitch command changes. While in ALT, will adjust altitude at a rate of 500 fpm. When not in ALT, will adjust pitch attitude at a rate of 0.7 deg/sec. Will cancel GS couple. The aircraft must pass through the glide slope again to allow GS recouple.

O. Yaw Damper Switch May be used to engage the yaw damper independent of the autopilot.

P. Control Wheel Steering Button When depressed, allows the pilot to manually control the aircraft (disengages the pitch and roll servos) without cancellation of any of the selected modes. Will engage the flight director if not previously engaged. Automatically synchronises the flight director/autopilot to the pitch attitude present when the CWS switch is released or to the present pressure altitude when in ALT hold mode. Will cancel GS couple. The aircraft must pass through the glide slope again to allow GS recouple.

Q. Manual Electric Trim Control Switches The left half of the switch provides power to engage the trim servo clutch. The right half controls the direction of motion of the trim servomotor. In order to operate the electric trim, both switches must be selected simultaneously. When the autopilot is engaged, operation of the electric trim will automatically disconnect the autopilot.

R. Autopilot Disconnect/Trim Interrupt (A/P Disc/Trim Inter) Switch When depressed will disengage the autopilot and yaw damper and cancel all operating flight director modes. When depressed and held, will do all above and also interrupt all trim power (stops trim motion), disengage the autopilot and yaw damper and cancel all operating flight director modes.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AUTOFLIGHT SYSTEM AND AVIONICS BEFORE TAKEOFF RELIABILITY TESTS 1.

2.

3.

Gyros -- Allow 3 – 4 minutes for gyros to come up to speed.

AVIONICS POWER 1 Switch – ON PREFLIGHT TEST BUTTON -- PRESS and OBSERVE. a. All annunciator lights on. b. TRIM annunciator flashing. c. After approximately 5 seconds, all annunciator lights off except the AP light, which will flash approximately 12 times and then remain off.

If the TRIM warning light stays on, the auto trim did not pass the pre-flight test. The autopilot circuit breaker should be pulled. Autopilot and manual electric trim will be inoperative.

4.

Manual Electric Trim -- TEST as follows : a. Actuate the left side split switch unit to the fore and aft positions. The trim wheel should not move. Rotate the trim wheel manually against the engaged clutch to check for override capability. b. Actuate the right-side split switch unit to the fore and aft positions. The trim wheel should not move and normal trim wheel force is required to move it manually. c. Press the A/P DISC/TRIM INTER switch down and hold. The electric trim should not operate.

5.

6.

7.

8.

9.

Flight Director -- ENGAGE by pressing the FD or CWS button.

Autopilot -- ENGAGE by pressing the AP ENG button.

Flight Controls -- MOVE fore, aft, left and right to verify that the autopilot can be overpowered.

A/P DISC/TRIM Button -- PRESS. Verify that the autopilot disconnects and all flight director modes are cancelled.

TRIM -- SET for take-off.

AUTOPILOT LIMITATIONS • • • • • •

During autopilot operation, a pilot with seat belt fastened must be seated in the left seat.

The autopilot and yaw damper must be OFF during take-off and landing.

The system is approved for Cat I ILS conditions only -- Approach Mode Selected.

The autopilot must be disconnected at 200 feet AGL. The autopilot must be OFF during use of standby flap system.

In accordance with FAA recommendation, use of basic pitch attitude hold mode is recommended during operation in severe turbulence.

MAXIMUM ALTITUDE LOSS DUE TO MALFUNCTION Cruise, Climb and Descent

500 ft.

Manoeuvring

100 ft.

100 ft.

Approach

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AUTOFLIGHT SYSTEM AND AVIONICS ATTITUDE INDICATOR COMMAND BAR Displays computed steering commands referenced to the symbolic aircraft. The command bar is visible only when the FD is selected. The command bar is visible only when the FD mode is selected.

The command bar will be biased out of view whenever the system is invalid or a flight director mode is not selected.

FCI SYMBOLIC AIRCRAFT The aircraft pitch and roll attitude is displayed by the relationship between the fixed symbolic aircraft and the moveable background. During flight director operation, the symbolic aircraft is flown to align it with the command bar to satisfy the flight director commands.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AUTOFLIGHT SYSTEM AND AVIONICS HORISONTAL SITUATION INDICATOR The HSI provides a pictorial presentation of the Aircraft deviation relative to VOR radials or localizer beams. It also displays glide slope deviations and gives heading reference with respect to magnetic north. The gyro is driven electrically.

NAV FLAG The flag is in view when the NAV receiver signal in inadequate. When a NAV flag is present in the HSI, the autopilot operation is not affected. The pilot must monitor the navigation indicator for a NAV flag to ensure that the autopilot and/or flight director are tracking valid navigation information.

LUBBER LINE Indicates aircraft magnetic heading on the compass card.

HEADING WARNING FLAG (HDG) When the flag is in view, heading display is invalid. If a HDG flag appears and a lateral mode (HDG, NAV or APP) is selected, the autopilot will be disengaged. The autopilot may be re-engaged in the basic wings-level mode along with any vertical mode. The CWS switch would be used to manually maneuver the aircraft laterally.

COURSE BEARING PONTER Indicates the selected VOR course or localizer course on the compass card. The selected VOR radial or localizer course remains set on the compass card when the compass card rotates.

TO/FROM FLAG INDICATOR Indicates direction of the VOR station relative to the selected course.

DUAL GLIDE SLOPE POINTERS Indicates the glide slope scale aircraft displacement from the glide slope beam center. Glide slope pointers in view indicate a usable glide slope signal is being received.

GLIDE SLOPE SCALES Indicates displacement from the glide slope beam center. A glide slope deviation bar displacement of 2 dots represents a full-scale deflection (0,7 degree) deviation above or below glide slope beam centreline.

COURSE DEVIATION BAR The center position of the omni bearing pointer moves laterally to pictorially indicate the relationship of the aircraft to the selected course. It indicates degrees of angular displacement from VOR radials or displacement in nautical miles from RNAV courses.

COURSE DEVIATION SCALE A course deviation bar displacement of 5 dots represents full scale (VOR = 10 degrees, LOC = 2½ degrees, RNAV = 5 nm) HEADING BUG Moved by the heading selector knob to select the desired heading.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AUTOFLIGHT SYSTEM AND AVIONICS SLAVING AND ACCESSORY AND COMPENSATOR UNIT This unit controls the compass system.

MANUAL/AUTOMATIC (FREE/SLAVE) COMPASS SLAVE SWITCH Selects either the manual or automatic slaving mode for the compass system.

CW/CCW COMPASS MANUAL SLAVE SWITCH With the switch in the FREE position, allows manual compass card slaving in clockwise or counter clockwise position. The switch is spring loaded to the center position.

SLAVING METER Indicates the difference between the displayed heading and magnetic heading. Up deflection indicates a clockwise error of the compass card.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AUTOFLIGHT SYSTEM AND AVIONICS WEATHER RADAR The Weather Radar System consists of a wing pod mounted receiver-transmitter and antenna and a panel mounted radar indicator.

OPERATING CONTROLS AND INDICATORS A. Function Selector Switch The function selector switch controls application of power and selects mode of operating for transmitting, testing, and warm-up. Switch positions are as follows : OFF

Primary power is removed from the system.

SBY

(Standby) After 30 seconds in this mode, places the system in operational ready status. Use during warm-up and in-flight periods when the system is not in use. The word STBY is displayed in the lower left corner.

TST

(Test) Selects test function to determine the operational ready status. A test pattern is displayed. No transmission is possible. The word TEST is displayed in the lower left corner.

ON

Selects the position for normal operation. Commences radar transmission. Defaults to WX mode and 80 nm range. WX will be displayed in the lower left corner and 80 will be displayed just above the right end of the top concentric range mark.

B. Antenna Tilt Control The tilt control adjusts the antenna to move the radar beam up to +15 degrees above the horizontal or to a maximum of -15 degrees below the horizontal position. The horizontal position is indicated as zero degrees on the control. The tilt angle selected is displayed in the upper right corner of the indicator.

C. Stabilisation (Stab) Pushbutton When pushed in, stabilisation is disabled. The words “STAB OFF” will flash on and off in the upper left corner. When in our position, the antenna stabilisation is restored. This feature keeps the antenna azimuth parallel to the ground, independent of aircraft attitude. The antenna compensates for up to ± 25° of aircraft pitch and roll.

D. Gain Control Permits adjusting the radar receiver gain in the terrain MAP mode only. In the test function as well as in all weather modes the receiver gain is pre-set.

E. Map (Ground Mapping) Pushbutton Selects ground mapping when pressed. When MAP mode is selected, the word MAP is displayed in the lower left corner. GAIN control becomes an operator function. Manual GAIN control is important in obtaining a definitive presentation during varying topographic conditions.

Prominent terrain features are presented in green; yellow and red. Magenta is not used.

F. WxA (WEATHER ALERT) PUSHBUTTON Selects weather alert mode when selected. WxA is displayed in the lower left corner of the screen.

Storm cells are displayed in up to four colours depending on level of intensity. Green, yellow, red and magenta. Magenta area of storm flashes between magenta and black to indicate strong storm cell.

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AUTOFLIGHT SYSTEM AND AVIONICS G. Wx (WEAHER) PUSHBUTTON Selects weather mode when selected. Wx is displayed in the lower left corner of the screen.

Operation for the Wx mode is the same as the WxA except the areas of strong rainfall appear as a steady magenta colour and will not flash between magenta and black as it does in WxA mode.

SAFETY PRECAUTIONS • •

Do not turn on or operate within 15 feet of ground personnel or metal objects.

Do not turn on or operate during refuelling operations.

MARGIN OF AVOIDANCE • • • • •

Avoid any cell by 5 nm if the temperature is above freezing and by 10 nm if below freezing.

When penetrating an area of storms and passing between two cells, a minimum corridor of 10 to 20 nm is required.

If the radar return is very intense with hooks and scallops, increase this distance accordingly.

Avoid by 10 nm or any cell that is changing shape rapidly.

Never fly under an overhang from a mature cumulonimbus cloud.

EXTENDING LIFE OF THE MAGNETRON • •

The system is designed so that full operation is possible 30 seconds after turn on.

The pilot may choose to leave the function switch in OFF rather than SBY if no significant weather is in the immediate area.

TURN-OFF PROCEDURE • •

Always turn the indicator to SBY or OFF before turning the avionics master switch off.

The system will power-down in about 5 seconds after switched to the OFF position, to allow time for the antenna to move to the down position.

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OPERATION DEPARTMENT

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AUTOFLIGHT SYSTEM AND AVIONICS

LIMITATIONS AIRSPEED MAXIMUM OPERATING SPEED (Vmo)

175 KIAS

MANOEUVRING SPEED (Va) 9062 LBS

143 KIAS

7500 LBS

137 KIAS

6250 LBS

125 KIAS

5000 LBS

112 KIAS

MAXIMUM FLAP EXTENDED SPEED (Vfe) O° TO 10° FLAPS

175 KIAS

10° TO 20° FLAPS

150 KIAS

20° TO 30° FLAPS

125 KIAS

ENGINE START CYCLE USING THE AIRPLANE BATTERY

30 seconds ON - 60 seconds OFF. 30 seconds ON - 60 seconds OFF. 30 seconds ON - 30 minutes OFF.

USING EXTERNAL POWER

20 seconds ON - 120 seconds OFF. 20 seconds ON - 120 seconds OFF. 20 seconds ON - 60 minutes OFF.

POWER PLANT (PT6A-114A) MINIMUM Ng FOR STARTING EMERGENCY AIRSTART Ng RISE IN ITT AND Ng FOR START

12% (preferably 18%).

Minimum 10%.

Maximum 10 seconds.

BATTERY START VOLTS

24 volts minimum.

AUXILIARY POWER START VOLTS

20 volts minimum.

AUXILIARY POWER UNIT CAPACITY

ISSUE 01 – REV. 00

800 to 1700 Amp.

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OPERATION DEPARTMENT

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AUTOFLIGHT SYSTEM AND AVIONICS

TORQUE FT-LBS

MAXIMUM ITT °C

GAS GENERATOR RPM % Ng

PROPELLER RPM

TAKEOFF

1865

805

101.6

1900

MAXIMUM CLIMB

1865

765

101.6

1900

MAXIMUM CRUISE

1865

740

101.6

1900

----

685

52 (MIN)

----

MAXIMUM REVERSE *

1865

805

101.6

1825

TRANSIENT **

2200

850

102.6

2090

----

1090

----

----

1865

805

101.6

1900

POWER SETTING

IDLE

STARTING MAXIMUM CONTINUOUS

* **

Reverse power operation is time limited to one minute.

These values are limited to 2 seconds.

OPERATING ALTITUDES NON-ICING CONDITIONS

25 000 ft.

ICING CONDITIONS CONDITIONS WITH ANY ICING ICE ON THE AIRFRAME

20 000 ft.

25 000 ft.

WEIGHTS MAXIMUM RAMP WEIGHT

9097 lbs.

MAXIMUM TAKEOFF WEIGHT

9062 lbs.

MAXIMUM LANDING WEIGHT

9000 lbs.

KNOWN ICING OPERATION

8550 lbs.

FORWARD CARGO HOLD

230 lbs.

CENTRE CARGO HOLD

310 lbs.

AFT CARGO HOLD

270 lbs.

AFT CARGO HOLD

280 lbs.

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OPERATION DEPARTMENT

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TRAINING HANDOUT PILOT CESSNA 208B GRAND CARAVAN AUTOFLIGHT SYSTEM AND AVIONICS FLIGHT LOAD FACTOR LIMITS FLAPS UP FLAPS DOWN

+3.8g, -1.52g. +2.4g.

FUEL LIMITATIONS TOTAL FUEL

2244 lbs

TOTAL UNUSABLE

20 lbs

TOTAL USABLE

2224 lbs

MAXIMUM IMBALANCE

200 lbs

MAX FULL RUDDER SIDESLIP DURATION 3 minutes (due to possible fuel starvation). With low fuel reserves (FUEL LOW annunciator ON), continuous uncoordinated flight with more than one quarter “ball” out of centre position is prohibited.

Unusable fuel quantity increases when more severe sideslip is maintained.

USE OF AVGAS The use of avgas is restricted to emergency use and shall not be used for more than 150 hours in one overhaul period. A mixture of one part avgas and three parts of Jet A, Jet A-1, JP-1 or JP-5 may be used for emergency purposes for a maximum of 450 hours per overhaul period. Avgas contains tetraethyl lead (TEL) to increase its critical temperature and pressure. When mixed with jet fuel, the result is that the lead ends up sticking to the turbine blades and vanes.

MINIMUM OIL QUANTITY Fill to 1½ quarts of MAX HOT or MAX COLD position on dipstick.

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SEPTEMBER 2021

OPERATION DEPARTMENT

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OWNERSHIP BY RGA – OPERATION DEPARTMENT