SOLENTA AVIATION AIRCRAFT TECHNICAL NOTES CESSNA 208B GRAND CARAVAN TABLE OF CONTENTS Cessna C208B Grand Caravan General Airplane Cockpit And Cabin Warning Systems Turbine Engine Theory PT6A-114A Powerplant Propeller Fuel System Electrical System Air-conditioning And Ventilation Oxygen Ice And Rain Protection Instruments Avionics Limitations Section 1 Section 2 Section 3 Section 4 Section 5 Section 6 Section 7 Section 8 Section 9 Section 10 Section 11 Section 12 Section 13 Section 14 © 2005 Solenta Aviation (Pty) Ltd- All Rights Reserved Issued 2006 i GENERAL AIRPLANE Table of Contents Introduction ................................................. 1-1 Fuselage...................................................... 1-2 Wings .......................................................... 1-2 Empennage................................................. 1-3 Cargo Pod ................................................... 1-3 Dimensions ................................................. 1-4 Landing Gear ...............................................1-5 Brake System...............................................1-5 Minimum Turning Distance ..........................1-6 Flight Controls..............................................1-7 Control Locks ...............................................1-8 Wing Flap System........................................1-9 Introduction The Cessna C208B Grand Caravan is an un-pressurised, all-metal, high wing, single-engine aeroplane equipped with a fixed tricycle landing gear. A composite cargo pod is optional equipment Figure 1-1 Cessna 208B Grand Caravan Fuselage The fuselage length for the C208B is 41 feet 7 inches (12.67 m), compared to 37 feet 7 inches (11.46 m) for the shortbody C208. 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. © Solenta Aviation (Pty) Ltd 1-1 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan • • 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 post. 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 (15.88 m). 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. Figure 1-2 C208B Wing Construction 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. Empennage The height is 15 feet 5 inches (4.70 m). The empennage consists of a conventional vertical stabiliser, rudder, horizontal stabiliser and elevator. The vertical stabiliser consists of a forward and aft spar, sheet © Solenta Aviation (Pty) Ltd 1-2 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 wrap around skin panels which also form the leading edges. Figure 1-3 C208B Stabilizer Vortex Generators 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. Figure 1-4 C208B Cargo Pod The pod has a load-carrying capacity of 1090 pounds (494 kg’s). 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. (104 kg’s) • Centre compartment fwd 310 lbs. (140 kg’s) • Centre compartment aft 270 lbs. (122 kg’s) • Aft compartment 280 lbs. (127 kg’s) 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. © Solenta Aviation (Pty) Ltd 1-3 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Dimensions Figure 1-5 C208B 3-View Length ...............................41’ 7” .........12.67 m Wingspan ..........................52’ 1” .........15.88 m Height ............................... 15’ 5”..................4.70 m Wheelbase Length......... 13’ 3½”..................4.05 m Hartzell Propeller Ground Clearance Standard Nose-gear Fork........... 14 ¼” ..36 cm Nose Strut Fully Compressed ...... 5 ½” ..14 cm Extended Nose-gear Fork .......... 17 ¾” ..45 cm Nose Strut Fully Compressed ......8 7/8” ..22 cm McCauley Propeller Ground Clearance Standard Nose-gear Fork ...........11 ¼” .........29 cm Nose Strut Fully Compressed.......2 ½” ...........6 cm Extended Nose-gear Fork ..........14 ¾” .........37 cm Nose Strut Fully Compressed...... 5 7/8” .........15 cm © Solenta Aviation (Pty) Ltd 1-4 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Landing Gear The landing gear is a fixed-gear, tricycle type with steerable nose wheel and two main wheels. To ensure the stability of the aircraft on the ground while loading, a tail-stand may be temporarily attached to the rear tie-down point. Figure 1-6 C208B Landing Gear And Tail-stand 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. The main landing gear has a “tear away” feature which reduces damage to the fuselage structure in the event of an accident. Nose gear shock absorption is provided by an oil-filled oleo 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. To improve operation from unpaved surfaces, the standard nose gear fork and tyres can be replaced by a 3" extended nose gear fork and oversized tyres. Figure 1-7 C208B Standard And Extended Nose Gear Forks, Nose Gear Over Travel Indicator 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 braking 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. A © Solenta Aviation (Pty) Ltd 1-5 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan single park brake handle is located on the main instrument panel, below the pilot’s control column. To set the park brake, apply brake pressure and pull the handle aft. To release the parking brake, push the handle fully in. Figure 1-8 C208B Park Brake Handle And Hydraulic Reservoir 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. Minimum Turning Distance Figure 1-9 C208B Minimum Turning Radius Minimum Turn Wheel Travel ..........28’ 1” .... 8.56 m Minimum Turn Wingtip Travel ........65’ 5” .. 19.94 m © Solenta Aviation (Pty) Ltd 1-6 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 shortbody models C208 and C208A, enhance lateral stability. These fences are also included as options on the longbody C208B where they are called “payload extenders” and allow an increase of 300lbs in the MTOW. Figure 1-10 C208B Wing Fence, Aileron and Trim Control • 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°. Figure 1-11 C208B Spoiler © Solenta Aviation (Pty) Ltd 1-7 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan • 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. Figure 1-12 C208B Elevator Trim Controls • 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. Figure 1-13 C208B Rudder Trim Control Trimming is not accomplished through a trim tab, but rather internally through the nose wheel steering bungee mechanically connected to the rudder control system. The trimming system acts against the steering bungee to displace the pedals and move the rudder. The rudder trim setting is controlled by a horizontal trim control wheel mounted on the control pedestal. Control Locks A manual control lock inserted into the pilot’s control column secures the ailerons in the neutral position and the elevators in slightly trailing-edge-down position. Ensure that the warning flag is over the sidewall switch panel. Figure 1-14 C208B Control Lock And Rudder Lock © Solenta Aviation (Pty) Ltd 1-8 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 onefourth 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 to prevent flap vibration. 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. Figure 1-15 C208B Flaps And Flap Controls • 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 © Solenta Aviation (Pty) Ltd 1-9 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. Figure 1-16 C208B Standby Flap Controls © Solenta Aviation (Pty) Ltd 1-10 Dec-2005 COCKPIT AND CABIN Table of Contents Instrument Panel ......................................... 2-1 Control Pedestal.......................................... 2-2 Left Sidewall Switch & C/B Panel ............... 2-2 Overhead Panel .......................................... 2-3 Pilot Seats ................................................... 2-3 Aircraft Doors .............................................. 2-3 Crew Entry Doors.........................................2-4 Passenger Door ...........................................2-4 Cargo Door ..................................................2-5 Interior Lighting ............................................2-6 Exterior Lighting ...........................................2-7 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 gauge, 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. Figure 2-1 C208B Instrument Panel 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. © Solenta Aviation (Pty) Ltd 2-1 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 centre and just to the right of the instrument panel. Located above the avionics stacks in the top centre 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 gauge, 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. Air-conditioning controls, if installed, are also located here. Control Pedestal A control pedestal extending from the centre 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. Figure 2-2 C208B Control Pedestal Left Sidewall Switch And Circuit Breaker Panel Control switches for the engine, electrical systems and avionics systems are located on a separate panel mounted on the left cabin sidewall adjacent to the pilot. The circuit breakers are located on the side of the panel, running down to the cockpit floor. Figure 2-3 C208B Left Sidewall Panel © Solenta Aviation (Pty) Ltd 2-2 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Overhead Panel The left and right fuel tank selectors, standby flap motor switches, left and right vent air control knobs, and left and right pilots’ vent air outlets are located on the overhead panel. Provisions for an optional oxygen shutoff valve and pressure gauge are also provided. Figure 2-4 C208B Overhead Panel Pilot 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 Thandle under the centre 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. As a safety feature, the seat is designed to deform with a high g-load, to minimise injury to the pilot. Later model aircraft seats are equipped with armrests which can be moved to the side and raised to a position beside the seat back for stowage. Figure 2-5 C208B Pilot Seat 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. Aircraft Doors Entry to and exit from the aircraft is accomplished through an entry door on each side of the cockpit and on the Passenger version only, through a two-piece, airstair door on the right side. A cargo door on the left side can also be used for cabin entry. © Solenta Aviation (Pty) Ltd 2-3 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan x Crew Entry Doors The left crew entry door incorporates a conventional interior and exterior door handle, a lock override knob a key-operated door lock and an openable storm 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-centre 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 placarded direction to unlock or lock the door. Figure 2-6 C208B Crew Entry Doors The door is held open by clipping a small chain, attached to the bottom of the door, to a small ring on the cowling. As this chain cannot be reached by the pilots from their seats, ensure that the chain is disconnected before entering the aircraft for flight. To hold the door open while the engine cowlings are open, an attachment arm from the lower cowling can be extended out to the bottom of the door. To allow easy access to the cockpit, folding ladders are fitted to the floor beside each door. Great caution must be taken when using the right-hand crew door due to the close proximity of the engine exhaust. Even after the engine has been shut down, the exhaust will remain hot for some time, and serious burns are possible if the exhaust is inadvertently touched. The right-hand crew ladder will also be a burning hazard if it remains extended with the engine running. x Passenger 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. © Solenta Aviation (Pty) Ltd 2-4 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Figure 2-7 C208B Passenger Door 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-centre 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. x Cargo Door 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. Figure 2-8 C208B Cargo Door 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. © Solenta Aviation (Pty) Ltd 2-5 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan In an emergency do not attempt to exit the Super Cargomaster 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. Interior Lighting Figure 2-9 C208B Interior Lighting Controls Instrument Lighting Instrument and control panel lighting is provided by integral, flood and post lights. Four concentric-type 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. Cabin Light Four cabin lights are installed in the interior of the airplane. These lights are controlled either by a twoposition 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. No Smoking/Fasten Seatbelt 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. 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. © Solenta Aviation (Pty) Ltd 2-6 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. Exterior Lighting Figure 2-10 C208B Exterior Lighting Controls 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. Landing Lights Two landing lights are mounted in the outboard section of each wing leading edge. They provide illumination forward and downward during takeoff 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 takeoff 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. 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. 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. © Solenta Aviation (Pty) Ltd 2-7 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. © Solenta Aviation (Pty) Ltd 2-8 Dec-2005 WARNING SYSTEMS Table of Contents Annunciator Panel....................................... 3-1 Fuel Selectors OFF Warning System ......... 3-2 Stall Warning System.................................. 3-3 Overspeed Warning System........................3-3 Engine Fire Detection System .....................3-3 Annunciator Panel An annunciator panel to advise the pilots of specific conditions in the aircraft systems is mounted below the glareshield on the captain’s instrument 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. Figure 3-1 C208B Annunciator Panel 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 preset minimum dim level for the green lamps and the following amber lamps: x Left and right fuel low. x Standby electrical power inoperative. x Standby electrical power on. Annunciator Cause For Illumination An excessive temperature condition and/or possible engine fire has occurred in the engine compartment. Engine oil pressure at or below 40+2 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 pressure in the fuel manifold is below 4.75 psi. © Solenta Aviation (Pty) Ltd 3-1 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan The starter-generator is operating in the starter mode. Electrical power is being supplied to the engine ignition system. Electrical system bus voltage is below 24.5 volts, and power is being supplied from the battery. The vacuum system suction is less than approximately 3.0” Hg. The fuel level in the reservoir tank is approximately half-full or less. Fuel quantity in the left fuel tank is approximately 25 USg’s (95 ltr’s/ approximately 146 lb’s) or less. Fuel quantity in the right fuel tank is approximately 25 USg’s (95 ltr’s/ approximately 146 lb’s) or less. Electrical power is not available from the standby alternator. The selected inverter or has not been turned ON (if installed). One or both fuel selectors are OFF, the fuel selector warning circuit breaker is not set or the start control circuit breaker is not set. The upper passenger door (passenger version only) and/or the upper cargo door are not latched. NiCad battery only - the electrolyte temperature is critically high (160ºF/ 71ºC) NiCad battery only - the electrolyte temperature is high (140ºF/ 60ºC) Metal chips have been detected in either the reduction gearbox case or the accessory gearbox case. The standby alternator is supplying electrical power to the bus. Electrical power is being supplied to the windshield anti-ice power relay (if installed). Pressure in the de-ice boot system has reached approximately 15 psi. Preset Dim 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. Fuel Selectors OFF 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: A. BOTH SELECTORS OFF The red FUEL SELECT OFF annunciator illuminates. A single fuel select off warning horn is activated. © Solenta Aviation (Pty) Ltd 3-2 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan B. DURING ENGINE START WITH EITHER SELECTOR OFF The red FUEL SELECT OFF annunciator illuminates. Both fuel select off warning horns are activated. C. 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 select 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 the pilot. 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 o 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. 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 preset 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 flows 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 (218°C) at the firewall, 625°F (329°C) around the exhaust, or 450°F (232°C) at the 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. © Solenta Aviation (Pty) Ltd 3-3 Dec-2005 TURBINE ENGINE THEORY Table of Contents Introduction ................................................. 4-1 Air Inlet Duct................................................ 4-2 Compressor................................................. 4-3 Centrifugal Compressor .............................. 4-3 Axial-flow Compressor ................................ 4-3 Compressor Stall......................................... 4-4 Combustion Chamber..................................4-4 Turbine.........................................................4-4 Exhaust ........................................................4-5 Turbo-propeller Engine Power Conversion .4-5 Fixed-shaft Engine .......................................4-5 Free-turbine Engine .....................................4-5 Introduction Turbine engines are made up of the following sections: x Air inlet x Compressor x Combustion section x Turbine x Exhaust x Accessory section x Support systems (starter, lubrication, fuel and auxiliary components such as anti-icing, airconditioning and pressurisation) Figure 4-1 PT6A-114A Overview © Solenta Aviation (Pty) Ltd 4-1 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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: x Ambient air density x Airspeed of the aircraft x Rotational speed of the compressor Compressor Two main functions of the compressor are to supply compressed air to the burners, and to 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. x 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. Figure 4-2 Exploded View Of Typical Single-stage Centrifugal Compressor © Solenta Aviation (Pty) Ltd 4-2 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Advantages of centrifugal compressors: x High pressure rise per stage. Up to 10:1 in a single stage. x Good efficiency (compression) over a wide rotational speed range, idle to full power. x Simplicity of manufacture and low cost. x Low weight. x Low starting power requirements. x More resistant to foreign object damage. Disadvantages of centrifugal compressors: x Large frontal area for a given airflow. x More than two stages is not practical due to energy loss in the airflow when making the turn from one impeller to the next. x 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. Figure 4-3 Typical 4-Stage Axial-flow Compressor 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 preset at a specific angle and are contoured similar to small propeller blades. Stators, which are stationary bladetype airfoils, 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: x High peak efficiencies (compressor pressures ratio) created by its straight through design allowing higher ram efficiency. © Solenta Aviation (Pty) Ltd 4-3 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan x x 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: x Difficulty and high cost of manufacture. x Relatively high weight. x High starting power requirements. x Low pressure rise per stage. x Good efficiency only over a narrow rotational speed range. x Very susceptible to foreign object damage. x Compressor Stall 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: x Anything that alters the proper operation of the compressor blades such as blade failure, foreign object damage, etc. x A fuel mixture that is too lean. x Abrupt pitching movement. x Excess fuel flow. x Engine operating speeds well above or below normal recommended operating speed. 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: x Mix the fuel and air in the proper manner for the ambient conditions to assure the best possible combustion. x Cool the hot gasses to a temperature within the normal operating range of the turbine wheel. x 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. © Solenta Aviation (Pty) Ltd 4-4 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. 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. Turbo-propeller Engine Power Conversion There are two main power output turbine designs: the fixed shaft design and the free turbine engine. x Fixed-shaft Engine 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 constant engine RPM, controlled descent capability due to prevention of windmilling overspeed, and rapid reverse thrust availability. x Free-turbine Engine 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: x Better control of propeller speed. x The propeller can be held at very low rpm during taxiing, with low noise and low blade erosion. x The engine is easier to start, especially in cold weather. x The propeller and the gearbox do not directly transmit vibrations into the gas generator. x A propeller brake can be installed during turnarounds. © Solenta Aviation (Pty) Ltd 4-5 Dec-2005 PT6A-114A POWERPLANT Table of Contents PT6A-114A Powerplant............................... 5-1 Free-turbine Reverse-flow Operation.......... 5-2 Engine Description And Operation.............. 5-2 Inertial Separator......................................... 5-5 Compressor Bleed Valve ............................ 5-5 Ignition System............................................ 5-6 Starting System........................................... 5-6 Engine Start Considerations ....................... 5-7 Exhaust ....................................................... 5-8 Engine Accessory Section ...........................5-8 Magnetic Chip Detectors .............................5-9 Engine Fuel System.....................................5-9 Lubrication System ....................................5-13 Engine Controls .........................................5-15 Engine Instruments ....................................5-16 Engine Limitations......................................5-18 Starter Limitations ......................................5-19 PT6A-114A Powerplant The Pratt & Whitney Canada PT6A-114A is a free-turbine engine with two independent turbines. The engine utilises two independent turbine sections: one driving the compressor in the gas generator section, and the second driving the propeller shaft through a reduction gearbox. The compressor section utilises 3 axial stages and 1 centrifugal stage to provide compressed air for the combustion chamber. Figure 5-1 PT6A-114A Engine, Internal Arrangement © Solenta Aviation (Pty) Ltd 5-1 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan The gas generator speed is 37,500 rpm = 100% NG. Maximum permissible speed is 38,100 RPM = 101.6% NG. The power turbine speed (NF) is 33,000 RPM at a propeller shaft speed of 1900 RPM. 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 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. Figure 5-2 C208B Engine Offset Down And Right Free-turbine Reverse-flow Operation The free-turbine design of the PT6A engine refers to the fact that the turbine sections rotate freely, having no physical connection between them. The compressor turbine drives the engine compressor and engine accessories. Dual power turbines drive the power section and propeller through the planetary reduction gearbox. The compressor and power turbines are mounted on separate shafts and are driven in opposite directions by the gas flow across them. Figure 5-3 Free-turbine Reverse-flow Turboprop Engine Reverse flow refers to the direction of airflow through the engine. Inlet air enters the compressor at the aft end of the engine, moves forward through the compressor section and the turbines, and is exhausted at the front of the engine. Engine Description And Operation 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 before entering through a © Solenta Aviation (Pty) Ltd 5-2 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan circular plenum chamber where it is directed to the compressor guide vanes. The compressor air inlet incorporates a mesh screen which will prevent entry of large particles, but does not filter the air. Inlet air is then ducted through an annular plenum chamber, formed by the compressor inlet case, where it is directed to the compressor. The compressor consists of three axial stages combined with a single centrifugal stage and assembled as an integral unit. A row of stator vanes, located between each stage of the compressor, diffuse the air, raise static air pressure and direct the air to the next stage of compression. The compressed air passes through diffuser tubes which turn the air through 90º in direction, and converts the velocity to static pressure. The diffused air then passes through straightening vanes to the annulus surrounding the combustion chamber liner assembly. The combustion chamber liner consists of two annular sections bolted together at the front, domeshaped end. The outer wrapper incorporates an integral large exit duct. The liner assembly has perforations of various sizes that allow entry of compressor delivery air. The flow of air changes direction 180º as it enters and mixes with fuel. The fuel/air mixture is ignited and the resultant expanding gases are directed to the turbines. The location of the inner liner eliminates the need for a long shaft between the compressor and compressor turbine, reducing the overall length and weight of the engine. Figure 5-5 PT6A Combustion Chamber Fuel is injected into the combustion liner through 14 simplex nozzles arranged in two sets of seven for ease of starting. Each nozzle is supplied by a dual manifold consisting of primary and secondary transfer tubes and adapters. The fuel/air mixture is ignited by two spark igniters which protrude into the liner. The resultant gases expand from the liner, reverse direction in the exit duct zone and pass through the compressor turbine inlet guide vanes to the compressor turbine. The guide vanes ensure that the expanding gases strike the turbine blades at the correct angle, with minimum loss of energy. The still expanding gases are then directed forward to the power turbine section. The power turbine drives the propeller shaft via a reduction gearbox. The compressor and power turbines are located in the approximate centre of the engine, with their respective shafts extending in opposite directions. This feature provides for simplified installation and maintenance inspection procedures. The exhaust gas from the power turbine is collected and ducted into a single exhaust duct assembly and directed to atmosphere on the right side of the aircraft. © Solenta Aviation (Pty) Ltd 5-3 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Figure 5-4 PT6A-114A Engine Cross Section © Solenta Aviation (Pty) Ltd 5-4 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Inter turbine temperature, T5, is monitored by an integral bus bar, probe and harness assembly installed between the compressor and power turbines, with the probes projecting into the gas path. A terminal block mounted on the gas generator case provides a connection point to the cockpit instrumentation. The engine oil supply is contained in an integral oil tank, which forms the rear section of the compressor inlet case. The tank has a capacity of 2.5 US gallons, 9.5 litres, and is provided with manual filler access and a direct oil level sight gauge. Fuel supplied to the engine is pressurised by an engine-driven fuel pump, and its flow to the fuel manifold is controlled by the FCU. The power turbine drives a propeller through a two-stage planetary reduction gearbox, located at the front of the engine. The gearbox contains an integral torque meter device, which is instrumented to provide an accurate indication of engine power. A magnetic chip detector is installed at the bottom of the gearbox. Figure 5-6 C208B Engine System Annunciators Inertial Separator 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 airfoil 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. Figure 5-7 C208B Inertial Separator Airflow Inertial separator operation is controlled by a T-handle located on the lower instrument panel. The handle is labelled BYPASS-PULL, NORMAL-PUSH. For details of the inertial separator operation refer to Section 11 – Ice And Rain Protection Systems. Compressor Bleed Valve At low NG RPM, the compressor axial stages produce more compressed air than the centrifugal stage can use. A pneumatic piston compressor bleed valve compensates for excess air flow at low RPM by bleeding axial stage air, P2.5, to reduce back pressure on the axial stages. This pressure relief helps prevent axial stage compressor stall. At low NG speeds the compressor bleed valve is open. As power is increased the valve begins to progressively close. Above approximately 90% NG the bleed valve is closed. If the compressor bleed valve were to stick closed at low NG speeds, compressor stall could result from an attempt to accelerate the engine to higher power. If the valve were to stick open at high NG speeds, power output © Solenta Aviation (Pty) Ltd 5-5 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan would be considerably reduced. With the valve open, at a given NG RPM, ITT will increase slightly and torque will decrease. 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. Figure 5-8 C208B Ignition And Start Switches NORM 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: x Air starts without starter assist. x Operation on water or slush covered runways. x During flight in heavy rain. x Inadvertent icing encounters until the inertial bypass separator has been in bypass for 5 minutes. x Near fuel exhaustion as indicated by the RESERVOIR FUEL LOW light. 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. © Solenta Aviation (Pty) Ltd 5-6 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. Engine Start Considerations x 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. x 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. x 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. x 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. x 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. x 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. x A dry motoring, within starter limitations after confirming that all fuel drainage has stopped, ensures that no fuel is trapped before the next start. x 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. © Solenta Aviation (Pty) Ltd 5-7 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Exhaust 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 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 seconds with a right crosswind as the cargo pod may be damaged by the hot exhaust gases. Engine Accessory Section 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 tachometer produces electric current which is used in conjunction with the gas generator RPM indicator. Torquemeter 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. © Solenta Aviation (Pty) Ltd 5-8 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 6 – Propeller Systems. Propeller Overspeed Governor See Section 6 – Propeller Systems. Engine Reduction Gearing 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 secondstage reduction gear and then to the propeller shaft. Reduction ratio is from a maximum power turbine speed of 33,000 RPM down to a propeller speed of 1900 RPM. Oil Breather Drain Can Some aircraft have an oil breather can mounted on the right lower engine mount truss. This can collects 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. Magnetic 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. Engine Fuel System The engine fuel system for the PT6A-114A consists of the following basic components; oil-to-fuel heat exchanger, an engine-driven fuel pump, a fuel control unit (FCU), a flow divider and dump valve, a dual fuel manifold with 14 simplex nozzles, and two fuel drain valves. The system also includes the fuel topping governor, covered separately in Section 6 – Propeller Systems. The system provides fuel flow to satisfy the speed and power demands of the engine. x 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. © Solenta Aviation (Pty) Ltd 5-9 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Figure 5-9 C208B Simplified Engine Fuel System x 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. x Fuel Control Unit (FCU) 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 meters proper fuel amounts for all modes of engine operation. Flow rates are calibrated for starting, acceleration and maximum power. The FCU compares NG with power lever setting, and regulates fuel to the engine fuel nozzles. The FCU also senses compressor discharge pressure, P3, and compares it to NG to establish acceleration or deceleration fuel flow limits. A pre-set minimum flow orifice guarantees sufficient fuel flow at all operating altitudes to sustain engine operation at minimum power. 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 © Solenta Aviation (Pty) Ltd 5-10 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. Figure 5-10 C208B Simplified FCU 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 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 © Solenta Aviation (Pty) Ltd 5-11 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. 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. 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. 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. x FCU Operation 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. 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 begin 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. 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. © Solenta Aviation (Pty) Ltd 5-12 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. 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. 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. 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 underspeeding condition in the reverse thrust range by controlling propeller speed with the NF governing section of the propeller control. 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 underspeed condition. 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. Lubrication System The PT6A engine lubrication system functions primarily to cool and lubricate engine bearings and bushings. It also provides oil to the propeller governor and propeller reversing control system. The main oil tank houses a gear-type engine-driven pressure pump, an oil pressure regulator, a cold pressure relief valve and an oil filter. The engine oil tank, an integral part of the compressor inlet case, is located in front of the accessory gearbox. 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: x Oil from the number 1 bearing sump to accessory gearbox by gravity. x Oil from the number 2 bearing sump is pumped to the accessory gearbox by number 2 scavenge pump. x Oil from the number 3 and 4 bearing sump is pumped to the accessory gearbox by the power turbine bearing scavenge pump. x 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. © Solenta Aviation (Pty) Ltd 5-13 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Figure 5-11 PT6A-114A Lubrication System © Solenta Aviation (Pty) Ltd 5-14 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan x 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 qts 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. 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 will be damaged. Figure 5-11 Engine Controls 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 annunciated 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. © Solenta Aviation (Pty) Ltd 5-15 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 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. Engine Instruments Figure 5-12 C208B Engine Instruments © Solenta Aviation (Pty) Ltd 5-16 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Torque Indicator The torque indicator indicates the torque being produced by the engine. One some Super Cargomaster aircrafts 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 (interturbine 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 right-hand 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 Indicator 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 85 to 105 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. Oil Temperature Indicator 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. © Solenta Aviation (Pty) Ltd 5-17 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Engine Limitations Airplane and engine limits are described in the Limitations section of the AFM. These limitations have been approved by the FAA, and must be observed when operating the C208B. The following engine operating limits chart provides important limitations for all operating conditions. OPERATING CONDITIONS FOR C208B INSTALLATION MAXIMUM ITT °C NG RPM % NG (2) PROP. RPM OIL PRESS. PSIG (3) OIL TEMP. °C (7) 1865 (1) 1865 1970 (4) 1865 1970 (4) 805 (10) 101.6 1900 85 TO 105 10 TO 99 765 101.6 1900 85 TO 105 0 TO 99 740 101.6 1900 85 TO 105 0 TO 99 ---- ---- 685 52 ---- 40 -40 TO 99 675 1865 805 101.6 1825 85 TO 105 0 TO 99 ---- 2400 (6) 850 102.6 2090 ---- 0 TO 99 ---- ---- 1090 ---- ---- ---- 675 1865 805 101.6 1900 85 TO 105 POWER SETTING sHP (9) TAKEOFF 675 MAXIMUM CLIMB MAXIMUM CRUISE IDLE MAXIMUM REVERSE (5) TRANSIENT (11) STARTING (11) MAXIMUM CONTINUOUS (8) 675 675 TORQUE FT-LBS (Minimum) (Minimum) -40 (Minimum) 10 TO 99 1. Maximum allowable torque decreases with altitude and temperature, as per the Engine Torque For Takeoff figure of AFM Section 5 Figure 5-8. 2. For Every 10ºC below -30ºC OAT, reduce the maximum allowable NG by 2.2% 3. Normal Oil Pressure is 85 – 135 psi at NG speeds above 72% with oil temperature between 60º – 70ºC. Oil pressures below 85 psi are undesirable, and should be tolerated only for the completion of the flight, preferably at a reduced power setting. Oil pressures below normal should be reported as an engine discrepancy and should be corrected before the next flight. 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 the minimum power possible to sustain flight. 4. Propeller RPM must be set so as not to exceed 675 sHP with torque above 1865 ft-lbs. Full 675 sHP rating is available only at RPM setting of 1800 or greater. 5. Reverse power operation limited to 1 minute. 6. Time limited to 20 seconds. 7. For increased oil service life, an oil temperature between 74ºC – 80ºC is recommended. A minimum oil temperature of 55ºC is recommended for fuel heater operation at takeoff. 8. Use of this rating is intended for abnormal situations (e.g. extreme icing, windshear, etc). 9. The maximum allowable sHP is 675. Less than 675 sHP is available under certain temperature and altitude conditions, as reflected in the takeoff, climb and cruise performance charts. 10. When the ITT exceed 765ºC, this power setting is time limited to 5 minutes. 11. Time limited to 2 seconds. During engine start, temperature is the most critical limit. The ITT starting limit of 1090ºC is limited to 2 seconds. During any start, if the indicator needle approaches this limit, the start should be aborted before the needle approaches the red triangle. © Solenta Aviation (Pty) Ltd 5-18 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Starter Limitations Engine starters are time-limited during the starting cycle to prevent the possibility of starter damage due to overheating. The time limitations on the starter vary depending on whether the start is being attempted using aircraft battery power, or external ground power. Aircraft Battery Start Cycle ................................ 30 seconds ON - 60 seconds OFF, 30 seconds ON - 60 seconds OFF, 30 seconds ON - 30 MINUTES OFF. External Power Start Cycle ............................... 20 seconds ON - 120 seconds OFF, 20 seconds ON - 120 seconds OFF, 20 seconds ON - 60 MINUTES OFF. © Solenta Aviation (Pty) Ltd 5-19 Dec-2005 PROPELLER SYSTEMS Table of Contents Propeller System......................................... 6-1 Blade Angle................................................. 6-2 Propeller Governor...................................... 6-2 Propeller Governor Operation ..................... 6-2 Low Pitch Stop ............................................ 6-3 Beta And Reverse Control .......................... 6-4 Beta And Reverse Control Operation ..........6-4 Propeller Overspeed Governor....................6-5 Power Turbine (NF) Overspeed Governor ...6-5 Power Lever.................................................6-7 Propeller Control Lever................................6-7 Feathering....................................................6-7 Propeller System The Cessna 208B is equipped with a constant speed, full-feathering, reversible, single-action, governor controlled, three bladed propeller, mounted on the output shaft of the reduction gearbox. Since the engine is a free turbine, with no mechanical connection between compressor and power turbines, the propeller can rotate freely on the power shaft when the engine is shut down. Propeller bungees and engine intake plugs are provided to prevent windmilling at zero oil pressure when the airplane is parked. The propeller is either a McCauley aluminium propeller or a Hartzell composite propeller. Figure 6-1 McCauley Prop Figure 6-2 Hartzell Prop Propeller pitch and speed are controlled by engine oil pressure supplied to the propeller dome through engine-driven propeller governors. A governor oil pump boosts oil pressure delivered by the engine oil system to a pressure high enough to control movement of the propeller blades. When oil pressure is present in the propeller dome, propeller pitch (blade angle) is controlled normally by the propeller governor or by the beta valve, depending upon the propeller’s mode of operation. As oil pressure increases, the propeller moves toward low pitch (high RPM). Loss of oil pressure will cause centrifugal counterweights and feathering springs to move propeller blades toward high pitch (low RPM). As oil pressure decreases during engine shutdown, the propeller automatically moves toward feather. © Solenta Aviation (Pty) Ltd 6-1 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan The minimum low pitch propeller position is determined by a mechanically-actuated hydraulic stop, referred to as the primary low pitch stop. The power lever controls beta and reverse blade angles by adjusting the low pitch stop position in beta and reverse ranges. Two governors (a propeller governor and a propeller overspeed governor) control propeller RPM. The primary governor controls the propeller through its normal governing range. The propeller control lever selects propeller RPM by adjusting the governor condition. Should the primary governor malfunction, the propeller overspeed governor prevents propeller speed from exceeding approximately 2000 RPM. The power turbine (NF) overspeed governor acts as a backup governor limiting propeller speed to 106% of that selected by the propeller lever. In the reverse range, the power turbine (NF) overspeed governor is reset, limiting propeller RPM to approximately 95% of the governor setting. The power turbine (NF) overspeed governor limits propeller RPM by reducing fuel flow to the engine. Blade Angle Blade angle is the angle between the chord of the propeller and the propeller’s plane of rotation. Because of the normal twist of the propeller, blade angle is different near the hub than it is near the tip. Blade angle for the C208B is measured at the chord, 42 inches from the propeller’s hub for a Hartzell Propeller, and at 30 inches for a McCauley propeller. Propeller Governor Figure 6-3 Blade Angles Also known as the Primary Governor on some other PT6 installations. 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. 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 centrebore 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 determine 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. The propeller governor can maintain any selected propeller speed from approximately 1600 RPM to 1900 RPM. Propeller Governor Operation The primary governor modulates oil pressure in the propeller dome to change blade angle to maintain a constant propeller speed. As oil pressure in the dome changes, propeller blade angles change to maintain the propeller speed the pilot has selected. For example, suppose an airplane is in normal cruising flight with the propellers set at 1750 RPM. If the pilot begins a descent without changing power, the airspeed will increase. This decreases the angle of attack of the propeller blades, causing less drag on the blades, thus causing the RPM to increase. The governor will sense this overspeed condition and increase blade angle to a higher pitch. The higher pitch increases the blade’s angle of attack, slowing it back to an on-speed 1750 RPM. © Solenta Aviation (Pty) Ltd 6-2 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Figure 6-4 Propeller Governor Operation Likewise if an airplane changes from cruise to climb attitude without a power change, the propeller RPM tends to decrease. The governor responds to this under-speed condition by decreasing blade angle to a lower pitch, and the RPM returns to its original value. Thus the governor is able to keep the variable pitch propeller at a constant speed. Power changes, as well as airspeed changes, cause the propeller to momentarily experience overspeed and under-speed conditions, but again the governor reacts to maintain the on-speed condition. Due to the smooth action of the governor, the pilot will notice few, if any, of these minor adjustments. There are times, however, when the primary governor is incapable of maintaining the selected RPM. For example, imagine an airplane approaching to land with its propellers set at 1900 RPM. As power and airspeed are both reduced, under-speed conditions exist which cause the governor to decrease blade angle as it attempts to restore the on-speed condition. If blade angle were allowed to decrease to its full reverse limit, aircraft control would be dramatically reduced as the propeller blades moved into a high drag disking position. To prevent this undesirable situation, a device is provided to stop the governor from selecting blade angles that are too low for safety. As blade angle is decreased by the governor, eventually the low pitch stop is reached. Blade angle then becomes fixed, preventing its continued movement toward a lower pitch. At the low pitch stop, the governor is prevented from restoring the on-speed condition, and propeller RPM decreases below the selected governor RPM setting. Once the low pitch stop is reached, blade angle cannot decrease further until the pilot selects beta or reverse. Low Pitch Stop With a non-reversing propeller the low pitch stop is simply at the low pitch limit of travel, determined by the propeller’s construction. But with a reversing propeller, extreme travel in the low pitch direction is past 0º, into reverse or negative blade angles. Consequently, the C208B’s propeller system has been designed to allow the low pitch stop to be repositioned when reversing is desired. The low pitch stop is created by mechanical linkage which senses blade angle. At the low pitch stop the linkage causes a valve to close, stopping the flow of oil into the propeller dome. Since more oil causes low pitch and reversing, blocking off oil flow creates a low pitch stop. The low pitch stop valve, commonly referred to as the “beta” valve, is spring-loaded to provide redundancy in the event of mechanical loss of beta valve control. Low pitch stop operation is determined by a mechanically monitored hydraulic stop. The propeller servo piston is connected by four spring-loaded sliding rods to the slip ring mounted behind the propeller. A carbon brush block riding on the slip ring transfers the movement of the slip ring through the propeller reversing lever to the beta valve of the governor. The initial forward motion of the beta valve blocks off the flow of oil to the propeller. Further motion forward dumps the oil from the propeller into the reduction gearbox sump. A mechanical stop limits the forward motion of the beta valve. Rearward motion of the beta valve does not affect normal propeller control. When the propeller is © Solenta Aviation (Pty) Ltd 6-3 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan rotating at a speed lower than that selected on the governor, the governor pump provides oil pressure to the servo piston and decreases the pitch of the propeller blades until the feedback of motion from the slip ring pulls the beta valve into a position blocking the supply of oil to the propeller, thus preventing further pitch changes. Beta and Reverse Control The position of the low pitch stop is controlled from the cockpit by the power lever. Whenever the power lever is at IDLE or above, this stop is set at 15.6º on McCauley propellers, and 9.0º on Hartzell propellers. Bringing the power lever aft of idle progressively repositions the low pitch stop to lower blade angles. 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 beta valve’s position. When the power lever is moved from idle to the reverse range, it repositions the beta valve to direct governor pressure to the propeller piston, decreasing blade angle through zero into a negative range. The travel of the propeller servo piston is fed back to the beta valve to null its position and, in effect, to provide infinite negative blade angles all the way to maximum reverse. The opposite will occur when the power lever is moved from full reverse to any forward position up to idle, thus providing the pilot with manual blade angle control for ground handling. The region of propeller travel between idle and ground fine is referred to as “beta”. In this range NG remains at idle. To enter the beta range, the pistol-trigger catch on the power lever must be lifted, and the lever brought aft past the idle stop. The aft stop of the beta range is called ground fine. With aft movement of the power lever, blade angle moves progressively from idle to ground fine. The region between ground fine and maximum reverse is referred to as the reverse range. In this range, NG progressively increases while the blade angle decreases. To enter the reverse range the power lever is moved aft of the beta range. Once the reverse range has been entered further aft movement of the power lever will cause the blade angle to progressively decrease from ground fine to maximum reverse. Beta and Reverse Control Operation Figure 6-5 Beta Range and Reverse Diagram © Solenta Aviation (Pty) Ltd 6-4 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan When the propeller blade angle approaches the low pitch stop setting, the four flanges extending from the dome make contact with four beta nuts. As propeller pitch angle continues to decrease each flange on the propeller dome pushes each beta nut and attached rod forward. As the rod moves forward it pulls the feedback ring forward. In turn a beta valve inside the governor is pulled into the oil cut-off position. The linkage is set to cut off oil supply to the dome when blade angle reaches 15.6º on McCauley propellers, and 9.0º on Hartzell propellers. This provides the governor with a hydraulic low pitch stop of for flight operations. If the low pitch stop were fixed at the propeller could not enter the beta and reverse range; however, the low pitch stop can be reset to allow the aircraft to enter into the beta and reverse ranges while the aircraft is on the ground. When the power lever is lifted back over the idle detent into the beta range, it is pulling back on the top of the reverse lever. As the reverse lever moves back, the beta valve is pushed back, reestablishing oil flow to the propeller dome. This allows propeller blade angle to go below the low pitch stop. As the propeller blades go below the low pitch stop, the propeller dome and feedback ring continue forward, eventually pulling the beta valve back into the oil cut-off position. Power Turbine (NF) Overspeed Governor Also known as the Fuel Topping Governor on some other PT6 installations. The primary propeller governor contains a power turbine (NF) overspeed governor which prevents power turbine overspeed if a propeller malfunctions. An overspeed could occur, for example, if a propeller blade were to stick in a fixed position during normal primary governor operation. In addition, during reverse thrust operation, the fuel topping governor is set below the speed selected by the primary governor to permit indirect control of propeller speed by the FCU servo system. The speed at which the power turbine (NF) overspeed governor operation occurs is determined by the speed selected with the propeller levers, and by the position of the reset lever. In the flight range the reset lever is set to regulate power turbine, N2, speed at approximately 106% of the propeller lever setting. In the ground range the lever is reset to 95% of the propeller lever position. 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. Fuel flow will be reduced, and engine power will decrease. When this occurs, propeller RPM normally remains constant, but it may decrease if propeller blades are frozen in a fixed position. Propeller Overspeed Governor Also known as the Overspeed Governor or Hydraulic Governor on some other PT6 installations. Since the PT6’s propeller is driven by a free turbine (independent of the engine’s compressor), overspeed can rapidly occur if the primary governor fails. The overspeed governor provides protection against excessive propeller speed in the event of primary governor malfunction. 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. © Solenta Aviation (Pty) Ltd 6-5 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Figure 6-6 Overspeed Governor Operation 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. Figure 6-7 Overspeed Governor Test Switch © Solenta Aviation (Pty) Ltd 6-6 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Power Lever The power lever is located on the power lever quadrant on the control pedestal. It is mechanically interconnected through a cam box to the FCU, reverse lever, beta valve and follow-up mechanism, and the propeller governor. The power lever quadrant permits movement of the power lever in the forward thrust (alpha) range from idle to maximum thrust, and in the beta/ reverse range from idle to maximum reverse. Mechanical stops in the power lever quadrant at the IDLE positions prevent inadvertent movement of the lever into the beta/ reverse range. To select beta or reverse, the pilot must lift up catches on the power lever to allow the power lever to be brought back past the stops. Figure 6-8 Control Pedestal In the forward thrust (alpha) range the power lever establishes NG by selecting a gas generator governor speed which results in a fuel flow that will maintain the selected NG RPM. In the beta range, the power lever controls the beta valve to reduce propeller blade angle, thus reducing residual propeller thrust. NG RPM is not affected in the beta range. In the reverse range, the power lever: • selects a blade angle proportionate to the aft travel of the lever • selects a fuel flow that will sustain the selected reverse power • resets the power turbine (NF) overspeed governor from its normal 106% to 95% of the propeller governor setting Therefore, RPM in reverse is a function of the primary propeller governor acting through the FCU to limit fuel flow and control propeller RPM in relation to power lever position. Propeller Control Lever Propeller RPM, within the primary governor range of 1600 to 1900 RPM, is set by the position of the propeller control lever. The full forward position sets the primary governor at 1900 RPM. In the full aft position, forward of the feathering detent, the primary governor is set at 1600 RPM. A detent at the low RPM position prevents inadvertent movement of the propeller lever into feather. Propeller Feathering The propeller can be manually feathered by moving the propeller lever full aft, past the detent, into feather. This action locks the governor’s pilot valve in the full up position, opens the feather valve, and all oil quickly drains from the propeller pitch mechanism. As oil is dumped from propeller servo chambers, the counterweights and springs drive the propeller blades to the feather position. Since the © Solenta Aviation (Pty) Ltd 6-7 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan propeller shaft and the NG shaft are not connected, the propeller can be feathered with the engine running; however to avoid excessive torque loads on the propeller gearbox, the engine should be at idle power when propellers are manually feathered. © Solenta Aviation (Pty) Ltd 6-8 Dec-2005 FUEL SYSTEM Table of Contents Fuel System Operation ............................... 7-1 Firewall Shutoff Valve ................................. 7-3 Fuel Tank Selectors .................................... 7-3 Auxiliary Boost Pump Switch ...................... 7-3 Fuel Flow Indicator...................................... 7-4 Fuel Quantity Indicators...............................7-4 Annunciators ................................................7-4 Fuel Drain Valves.........................................7-5 Fuel Drain (EPA) Can ..................................7-5 Fuel Pump Drain Reservoir .........................7-6 Fuel System Operation Two vented, integral fuel tanks with shutoff valves, a fuel selectors off warning system, a fuel reservoir, an ejector fuel pump, an electric auxiliary boost pump, a reservoir manifold assembly, a firewall shutoff valve, a fuel filter, an oil-to-fuel heater, an engine-driven fuel pump, a fuel control unit, a flow divider, dual manifolds, 14 fuel nozzle assemblies and a fuel drain can and drain. 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. © Solenta Aviation (Pty) Ltd 7-1 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Figure 7-1 C208B Fuel System Schematic 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. © Solenta Aviation (Pty) Ltd 7-2 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 centre which locks the knob in position when the button is released. Figure 7-2 C208B Firewall Shutoff Valve 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. Figure 7-3 C208B Fuel Tank Selectors Auxiliary Boost Pump Switch The auxiliary boost pump switch is located on the pilot’s side panel, and is labelled FUEL BOOST and has OFF, NORM and ON positions. 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. Figure 7-4 C208B Auxiliary Boost Pump Switch © Solenta Aviation (Pty) Ltd 7-3 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 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. Figure 7-5 C208B Fuel System Gauges 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 2½ 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 Figure 7-6 C208B Fuel System Annunciators 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 PRESS 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. 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. © Solenta Aviation (Pty) Ltd 7-4 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan LEFT And RIGHT FUEL LOW Two amber fuel low warning annunciators are labelled LEFT FUEL LOW and RIGHT FUEL LOW. Each annunciator will illuminate when the fuel in the respective tank is less than 25 gallons. (95 ltr’s/ approximately 146 lb’s) FUEL SELECT OFF A red fuel selectors off annunciator labelled FUEL SELECT OFF will illuminate when one or both fuel selectors are OFF, the fuel selector warning circuit breaker is not set or the start control circuit breaker is not set. 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. Figure 7-7 C208B Wing Fuel Tank, And Reservoir Tank Drains 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 (EPA) 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. Figure 7-8 C208B Fuel Filter, EPA Can And Fuel Pump Drain © Solenta Aviation (Pty) Ltd 7-5 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. © Solenta Aviation (Pty) Ltd 7-6 Dec-2005 ELECTRICAL SYSTEM Table of Contents Introduction ................................................. 8-1 Electrical Buses........................................... 8-2 Starter-generator......................................... 8-2 Generator Control Unit ................................ 8-2 Ground Power Monitor ................................ 8-2 Ground service Plug Receptacle ................ 8-3 External Power Switch ................................ 8-3 Battery ......................................................... 8-3 Battery Switch ..............................................8-4 Generator Switch .........................................8-4 Avionics Power Switches.............................8-4 Avionics Bus Tie Switch...............................8-4 Circuit Breakers ...........................................8-5 Volt/Ammeter ...............................................8-5 Standby Electrical System ...........................8-5 Introduction Figure 8-1 C208B Electrical Schematic © Solenta Aviation (Pty) Ltd 8-1 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. Figure 8-2 C208B Electrical System Annunciators 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 cutout 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 monitor 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. © Solenta Aviation (Pty) Ltd 8-2 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 or maintenance. Figure 8-3 C208B Ground Service Plug Receptacle 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 three-position switch. 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. Figure 8-4 C208B Left Sidewall Panel 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. 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. © Solenta Aviation (Pty) Ltd 8-3 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 Annunciator This amber annunciator illuminates when the NiCad battery temperature is 140 - 160°C. • Battery Overheat Annunciator This red annunciator illuminates when the NiCad battery temperature exceeds 160°C. Immediately turn the battery switch OFF. 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. 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 © Solenta Aviation (Pty) Ltd 8-4 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 beakers. 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 are 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. Figure 8-5 C208B Volt/Ammeter 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. 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. © Solenta Aviation (Pty) Ltd 8-5 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 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. © Solenta Aviation (Pty) Ltd 8-6 Dec-2005 AIR-CONDITIONING & VENTILATION SYSTEMS Table of Contents Ventilating Air Inlets .................................... 9-1 Ventilation Fans .......................................... 9-1 Instrument Panel Vent Knobs ..................... 9-2 Cabin Heating System ................................ 9-2 Heating Controls ..........................................9-3 Air-conditioning System ...............................9-5 Air-conditioning Controls..............................9-6 Air-conditioner Operation.............................9-6 Ventilating Air Inlets 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 Fans (Non-Air-conditioner Equipped Aircraft) An optional ventilation fan system may be installed to provide supplemental cabin ventilation. The system controls are located on the overhead console and consists of two rotary-type vent air controls, labelled VENT AIR. Figure 9-1 C208B Ventilation Fan Controls 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. © Solenta Aviation (Pty) Ltd 9-1 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. 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. Figure 9-2 C208B Cabin Air System, Including Ventilation Fans (For ease of display the Cabin Heater Outlets, and Cabin Overhead Ventilating Outlets have been removed from the diagram) © Solenta Aviation (Pty) Ltd 9-2 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Cabin Heating Controls The cabin heating controls are located at the bottom of the instrument panel, above the pedestal. Figure 9-3 C208B Cabin Heating Controls • 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. • Bleed Air Heat Switch The BLEED AIR HEAT switch is 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. • 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. 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. © Solenta Aviation (Pty) Ltd 9-3 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. 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. • 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 Super Cargomaster version, and on the cabin sidewalls at floor level for the Grand Caravan 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. Figure 9-4 Cabin Heat Firewall Shutoff Knob 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. © Solenta Aviation (Pty) Ltd 9-4 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Air-conditioning System An optional air-conditioning system may be installed into Grand Caravan version aircraft. The airconditioning system is able to provide cooling air both on the ground and in flight. A belt-driven compressor is located on the engine accessory section. There are three evaporator units with integral blowers, one in each wing root area, and one in the tail cone behind the aft bulkhead. The evaporator units direct cooled air to a series of overhead outlets in the cabin headliner. The system condenser is mounted beneath the engine, and is provided with a louvered air inlet and outlet on the lower left engine cowling. Refrigerant lines under the floorboards and in the fuselage sides connect the compressor, evaporators and the condenser. Figure 9-5 C208B Air-conditioning System When the air-conditioning system is operating, cooled air is supplied to the cabin through 16 overhead adjustable outlets, one above each pilot, 11 above the passengers, and 3 directing air forward from the rear cabin bulkhead. © Solenta Aviation (Pty) Ltd 9-5 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Air-conditioning Controls The cabin heating controls are located at the bottom of the instrument panel, above the pedestal, next to the cabin heat controls. Figure 9-6 C208B Air-conditioning Controls System electrical protection is provided by for 15 Amp circuit breakers, labelled AIR COND CONT, LEFT VENT BLWR, RIGHT VENT BLWR, and AFT VENT BLWR. • Air-conditioning Switch A three-position master switch controls the air-conditioning system. Placing the switch in the COOL position starts the system compressor and evaporator fans. With the switch in the VENTILATE position, power is only provided to the system evaporator fans, providing uncooled ventilating air to the cabin. • Air-conditioning Fan Switches The 3 two-position AC Fan switches provide separate HIGH or LOW speed control of each evaporator fan. Air-conditioner Operation As the air-conditioning compressor is belt-driven through the engine accessory section, significant engine power loss may be evident with the engine at low power on the ground, and at take-off power. Under very hot OAT, or high ground elevation conditions, the idle ITT may exceed the maximum idle limitation of 685ºC. In this case, advance the fuel condition lever towards HIGH IDLE to increase the NG as required to maintain the ITT within limits. For increased cooling while on the ground, select the fan switches to HIGH, and increase NG to 6065% for a higher air-conditioning compressor RPM. Once the cabin has reached a comfortable temperature the fan switches can be repositioned to LOW. Ground operation of the air-conditioner with the propeller in beta range for prolonged periods will cause the air-conditioning compressor pressure safety switch to disengage the compressor clutch, and therefore should be avoided. If the temperature of the air coming from the outlets does not start to cool within a minute or two, the system may be malfunctioning, and should be turned off. Due to the loss of engine power at take-off, it is recommended to place the air-conditioning master switch to VENTILATE immediately before take-off, and to return it to COOL only once safely airborne if air-conditioning is still required in flight. Likewise, before landing the air-conditioning master switch should be placed to VENTILATE to ensure maximum power is available in the event of a go-around. © Solenta Aviation (Pty) Ltd 9-6 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan During extended flight in high temperatures and humidity, the evaporator coils may frost over. Normally the compressor cycles off when temperatures in the evaporators near freezing. If frost does form, as evidenced by reduced cooling airflow, turn the air-conditioning master switch to VENTILATE and select the fan switches to HIGH. This should increase evaporator discharge temperature sufficiently to clear the frost. There is a 10–25 fpm reduction in aircraft climb performance, 1-2 KTAS decrease in cruise performance, and approximately 1% increase in fuel required for a given flight with the air-conditioner operational. © Solenta Aviation (Pty) Ltd 9-7 Dec-2005 OXYGEN SYSTEM Table of Contents Oxygen System......................................... 10-1 Oxygen System 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. Figure 10-1 C208B Oxygen Controls 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. © Solenta Aviation (Pty) Ltd 10-1 Dec-2005 ICE & RAIN PROTECTION SYSTEMS Table of Contents Introduction ............................................... 11-1 Pneumatic De-ice Boots............................ 11-1 Propeller Anti-ice Boots............................. 11-2 Windshield Anti-ice Panel ......................... 11-2 Pitot And Stall Warning Heat .....................11-3 Ice Detector Light.......................................11-3 Engine Inertial Separator ...........................11-3 Introduction 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 Figure 11-1 C208B Ice Protection Controls All ice protection, with the exception of the manually activated engine inertial separator, are electrically controlled. 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 de-ice boots. © Solenta Aviation (Pty) Ltd 11-1 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 leading edge 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. Figure 11-2 C208B Pneumatic De-ice Boots And Propeller De-ice 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. 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 © Solenta Aviation (Pty) Ltd 11-2 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. Figure 11-3 C208B Windshield Anti-ice Panel Pitot And Stall Warning Heat 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. The stall warning 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. Ice Detector Light An ice detector light is flush mounted near the upper left corner of the windshield, to illuminate the pilot’s wing. The switch is spring-loaded to the off position and must be held in the ON position to illuminate the light. Engine Inertial Separator The inertial separator 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. The handle is very stiff to move back into the NORMAL position due to the airloads on the moveable vanes. When using the inertial separator in icing conditions, ensure that the handle is only returned to the NORMAL position after engine shutdown, due to the fact that ice build-up can still be present inside the inertial separator even though external signs of icing have dissipated. © Solenta Aviation (Pty) Ltd 11-3 Dec-2005 INSTRUMENTS Table of Contents Pitot-Static System.................................... 12-1 Vacuum System ........................................ 12-2 Suction Gauge .......................................... 12-2 Vacuum Low Warning Annunciator ...........12-2 Electrically Operated Gyros .......................12-3 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. Figure 12-1 C208B Static Controls 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. © Solenta Aviation (Pty) Ltd 12-1 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. Figure 12-2 C208B Typical Vacuum System • Suction Gauge 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. © Solenta Aviation (Pty) Ltd 12-2 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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 takeoff 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 takeoff is 10 minutes or more, allow the gyro to run as caging is not necessary. © Solenta Aviation (Pty) Ltd 12-3 Dec-2005 AVIONICS Table of Contents King KFC-150 Autopilot System ............... 13-1 Definitions And Terms............................... 13-2 KC-192 Autopilot And FD Computer......... 13-2 KA-185 Remote Mode Annunciator .......... 13-4 Autopilot Control Yoke Switches............... 13-5 Before Take-off Reliability Tests ............... 13-5 Autopilot Limitations .................................. 13-6 Flight Command Indicator (FCI) ................13-6 Horizontal Situation Indicator (HSI) ...........13-7 Slaving Accessory And Compensator .......13-8 Weather Radar...........................................13-9 Radar Safety Precautions........................13-10 Operational Procedures...........................13-10 King 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 autotrim during autopilot operation and manual electric trim for the pilot. Trim faults are visually and aurally annunciated. A lockout device prevents autopilot engagement until the system has been successfully preflight 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. Figure 13-1 C208B KFC-150 Autopilot System Components © Solenta Aviation (Pty) Ltd 13-1 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Definitions And Terms 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. Flight Director (FD) 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 serves 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. KC-192 Autopilot And Flight Director Computer The autopilot system is controlled by the Autopilot and FD computer, located at the lower right side of the avionics stack. Figure 13-2 KC-192 Autopilot And Flight Director Computer • Controls And Indicators YD – Yaw Damper Annunciator Illuminates when the yaw damper (optional) is engaged. © Solenta Aviation (Pty) Ltd 13-2 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Mode Annunciators FD – Flight Director Mode ALT – Altitude Hold Mode HDG – Heading Mode NAV – Navigation Mode APR – Approach Mode BC – Back Course Mode Illuminate when a mode is selected by the corresponding mode selector button. GS – Glide Slope 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. TRIM – Trim Warning Light Illuminates continuously whenever trim power is not on or the system has not been preflight tested. The TRIM warning light illuminates and an audible warning will sound whenever a trim fault is detected. The autotrim 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. AP – Autopilot Annunciator Illuminates continuously when the autopilot is engaged. Flashes 12 times whenever the autopilot is disengaged (an aural alert will sound for 2 seconds). AP ENG – Autopilot Engage Button When pushed, engages the autopilot when all logic conditions are met. TEST – Preflight Test Button When momentarily pushed, initiates preflight test sequence which: • Automatically turns on all lights. • Tests the roll and pitch rate monitors. • Tests the autotrim fault monitor. • Checks the manual trim drive voltage. • Tests all autopilot valid and dump logic. • If the preflight 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 preflight tests are successfully passed. BC – Back Course Approach 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. APR –Approach 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. NAV – Navigation 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 © Solenta Aviation (Pty) Ltd 13-3 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan capture sequence is initiated. On the Remote Mode Annunciator, NAV ARM will annunciate until the automatic capture sequence is initiated. At beam capture, NAV CPLD will annunciate. HDG – Heading 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. ALT – Altitude Hold 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. FD – Flight Director 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. DN/UP – Vertical Trim Control A spring-loaded-to-centre 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. YAW DAMP – Yaw Damper Switch May be used to engage the optional yaw damper independently of the autopilot. KA-185 Remote Mode Annunciator The remote mode annunciator provides annunciators within the pilot’s normal instrument scan area, as well as three marker beacon lights. Figure 13-3 KA-185 Remote Mode Annunciator ARM – Armed Annunciator Illuminates continuously along with NAV or APR when either mode button is depressed. The ARM annunciator will remain illuminated until the automatic capture sequence is initiated, at which time ARM will extinguish and CPLD will illuminate. CPLD – Coupled Annunciator Illuminates continuously along with NAV or APR when either mode button is depressed. The ARM annunciator will remain illuminated until the automatic capture sequence is initiated, at which time ARM will extinguish and CPLD will illuminate. © Solenta Aviation (Pty) Ltd 13-4 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Autopilot Control Yoke Switches Autopilot and electric trim controls are provided only on the captain’s control yoke. Figure 13-4 Autopilot Control Yoke Switches A/P DISC/TRIM INTER – Autopilot Disconnect And Trim Interrupt 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. 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. CWS – 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. Before Take-off Reliability Tests 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 autotrim did not pass the preflight test. The autopilot circuit breaker should be pulled. Autopilot and manual electric trim will be inoperative. 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. © Solenta Aviation (Pty) Ltd 13-5 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan c. Press the A/P DISC/TRIM INTER switch down and hold. The electric trim should not operate. 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 takeoff 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 takeoff 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 Approach ................................................................. 100 ft Flight Command Indicator (FCI) 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. Figure 13-5 KI-256 Flight Command Indicator 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. © Solenta Aviation (Pty) Ltd 13-6 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Horizontal Situation Indicator (HSI) Figure 13-6 KI-525A Horizontal 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 Flap (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 manoeuvre the aircraft laterally. Course Bearing Pointer 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. © Solenta Aviation (Pty) Ltd 13-7 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Dual Scale Slope Pointers Indicates the glide slope scale aircraft displacement from the glide slope beam centre. Glide slope pointers in view indicate a usable glide slope signal is being received. Glide Slope Scales Indicates displacement from the glide slope beam centre. 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 centre 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. Slaving Accessory And Compensator This unit controls the compass system. Figure 13-7 KA-51B Slaving Accessory And Compensator Unit MAN/AUTO – 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 counterclockwise position. The switch is spring loaded to the centre position. Slaving Meter Indicates the difference between the displayed heading and magnetic heading. Up deflection indicates a clockwise error of the compass card. © Solenta Aviation (Pty) Ltd 13-8 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Weather Radar The Weather Radar System consists of a receiver-transmitter and antenna mounted in a pod on the right-hand wing, and a panel mounted radar indicator. There a several different radar installations for the C208B. The most popular radar installation is the Narco (Bendix/King) KWX-56, however the same basic details should also be applicable to other installations. Figure 13-8 Narco KWX-56 Radar Indicator, And Wing Pod Installation 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. TILT - 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. STAB – Stabilisation Pushbutton When pushed in, stabilisation is disabled. The words “STAB OFF” will flash on and off in the upper left corner. When in out 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. GAIN – 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 preset. © Solenta Aviation (Pty) Ltd 13-9 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan 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. 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. Wx – Weather 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. Radar 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. Operational Procedures • • • • 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. 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. • 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. © Solenta Aviation (Pty) Ltd 13-10 Dec-2005 LIMITATIONS Table of Contents Airspeeds .................................................. 14-1 Engine Start Cycles................................... 14-1 Power Plant............................................... 14-1 Operating Altitudes.................................... 14-2 Weights ..................................................... 14-2 Flight Load Factor Limits ...........................14-2 Maximum Full Rudder Sideslip Duration ...14-2 Fuel ............................................................14-3 Use Of Avgas.............................................14-3 Minimum Oil Quantity ................................14-3 Airspeeds Maximum Operating Speed (VMO) ..................175 KIAS Manoeuvring Speed (VA) 8,750 lbs. ....................................................................148 KIAS 7,500 lbs. ....................................................................137 KIAS 6,250 lbs. ....................................................................125 KIAS 5,000 lbs. ....................................................................112 KIAS Maximum Flap Extended Speed (VFE) 0° To 10° Flaps...........................................................175 KIAS 10° To 20° Flaps.........................................................150 KIAS 20° To 30° Flaps.........................................................125 KIAS Figure 14-1 C208B ASI Engine Start Cycles Aircraft Battery Start Cycle ................................ 30 seconds ON - 60 seconds OFF, 30 seconds ON - 60 seconds OFF, 30 seconds ON - 30 MINUTES OFF. External Power Start Cycle ............................... 20 seconds ON - 120 seconds OFF, 20 seconds ON - 120 seconds OFF, 20 seconds ON - 60 MINUTES OFF. Powerplant (PT6A-114A) Minimum NG For Starting ................................... 12% (preferably 18%) Emergency Airstart NG ....................................... Minimum 10% Rise In NG And ITT During Start ..................... Maximum 10 seconds Battery Start Minimum Voltage........................ 24 volts minimum © Solenta Aviation (Pty) Ltd 14-1 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan APU Start Minimum Voltage............................. 20 volts minimum APU Capacity ........................................................ 800 to 1700 Amp Figure 14-2 C208B Engine Instruments MAXIMUM ITT °C GAS GENERATOR RPM % NG PROPELLER RPM 805 101.6 1900 765 101.6 1900 1865 740 101.6 1900 ---- 685 52 (Minimum) ---- 1865 805 101.6 1825 2400 850 102.6 2090 ---- 1090 ---- ---- 1865 805 101.6 1900 POWER SETTING TORQUE FT-LBS TAKEOFF 1865 MAXIMUM CLIMB 1865 MAXIMUM CRUISE IDLE MAXIMUM REVERSE (1 Minute Maximum) TRANSIENT (2 Second Maximum) STARTING (2 Second Maximum) MAXIMUM CONTINUOUS (20 sec’s max.) (5 mins max.) Operating Altitudes Non-icing Conditions .....................................................25,000 ft Icing Conditions.............................................................20,000 ft Conditions With Any Ice On The Airframe ...................20,000 ft Weights Maximum Ramp Weight............................................... 8,785 lbs Maximum Take-off Weight........................................... 8,750 lbs Maximum Landing Weight ........................................... 8,500 lbs Known Icing Operation................................................. 8,550 lbs Forward Cargo Compartment......................................... 230 lbs Centre Forward Cargo Compartment............................. 310 lbs Centre Aft Cargo Compartment...................................... 270 lbs Aft Cargo Compartment .................................................. 280 lbs Flight Load Factor Limits Flaps UP ................................................................ +3.8g, -1.52g Flaps DOWN ......................................................................+2.4g © Solenta Aviation (Pty) Ltd 14-2 Dec-2005 AIRCRAFT TECHNICAL NOTES Cessna 208B Grand Caravan Maximum 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. Fuel Total Fuel...................................................................... 2,244 lbs Total Unusable Fuel .......................................................... 20 lbs Total Useable Fuel ....................................................... 2,224 lbs Maximum Fuel Imbalance............................................... 200 lbs 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. © Solenta Aviation (Pty) Ltd 14-3 Dec-2005
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