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North American T-28C Trojan V2 for X-Plane V9.70
Flight manual


The T-28B and T-28C are straight-wing, high-performance, two-place trainers equipped with dual controls. The T-28C is a carrier version of the T-28B and maintains the same outward appearance except for the arresting hook. A speed brake is located on the bottom of the fuselage and, with a few exceptions, both cockpits contain identical controls and instruments. The cockpits were designed and arranged to be as much like fighter cockpits as practical. For armament training, bombs may be carried externally under each wing panel. When any configuration of armament is carried, a fire control system is installed in the front cockpit.


Approximate overall dimensions are:

  • Length : 33 feet
  • Wing Span : 41 feet
  • Height (to top of fin) : 13 feet


Power is provided by a nine-cylinder, radial, air-cooled Wright Cyclone engine, Model R1820-86A. At take-off and military power, the engine develops 1425 horsepower. Engine exhaust outlets on each side of the cowl are designed to utilize the additional thrust available from the jet effect of the exhaust. The engine is equipped with a single-stage, two·speed, engine·driven supercharger, a direct-cranking starter, and an injection-type carburetor incorporating an electric primer valve.

North American T-28C Trojan for X-Plane by Khamsin & Arno54


Engine throttle, mixture, supercharger, and carburetor air controls are located on the left side of each cockpit and are interconnected between cockpits to move simultaneously. . Cylinder head temperature and oil temperature are controlled simultaneously by electrically actuated cowl and oil cooler flaps.


Each throttle is provided with a take-off stop in the quadrant so the pilot can feel when the throttle has been advanced to take-off power (at sea level). Pushing the throttle through the stop at sea level will cause the maximum allowable manifold pressure to be exceeded. At altitudes above sea level, the throttle may be pushed beyond the stop as long as the manifold pressure is kept below the maximum limit. Forward movement of the throttle actuates a mechanically linked carburetor accelerating pump. The throttle grip contains the SPEED BRAKE switch.


The mixture control lever has three positions: RICH, NORMAL, and IDLE CUTOFF. The RICH position is used for all ground operation, take-off, climb, descent, and landing; the NORMAL position is used for all other normal flight conditions. The IDLE CUTOFF position shuts off fuel flow at the carburetor to stop the engine. Mixture is adjustable between detents.

Note Fuel is injected into the impeller section of the engine if the boost pump is operating and the mixture control is not in the IDLE CUTOFF position, however, with sufficient airspeed, the engine may absorb fuel and self-detonate. Don't mistake it for a running engine, as this will provide almost no thrust.


Operating speed ratio of the two-speed supercharger is selected by the position of the supercharger control handle, located on the throttle quadrant in each cockpit. When the handle is at the LOW (up) position, the supercharger is set at low blower; when the handle is at the HIGH (down) position, the super-charger is set at high blower.


A carburetor air control is located below the throttle quadrant in each cockpit. With the handle at the DOWN-FORTH position, the ram-air duct is closed and heated air from the area aft of the engine is drawn into the carburetor("Carb Heat" is on). As the handle is moved UP and BACKWARD , heated air is mixed with cold ram air to obtain the desired carburetor air temperature.("Carb Heat" is off)


A standard magneto ignition switch is located on the right instrument subpanel in each cockpit. Switch positions are OFF, L, R, BOTH and IGNITION. The L and R positions are provided to individually check engine operation on the left or right ignition system.


The direct-cranking electric starter is controlled by a guarded push button on the right console in each cockpit. Holding the STARTER button down operates the starter. The starter can be powered by the battery when external power is not available; however, this procedure causes a heavy current drain from the battery and should be used only when external power is not available. The starter is powered by the primary d-c bus.


The engine fuel priming system for starting is controlled by a PRIMER button on the right forward console in each cockpit. Depressing this button opens an electric primer valve on the carburetor, permitting pressurized fuel from the carburetor to be injected into the engine blower section. Fuel pressure for priming is provided by the fuel booster pump. The system is very sensitive and care must be used to avoid overpriming. The primer system is powered by the primary d-c bus. Note : no effect is noticeable in Xplane, although the switch is operationnal.


Identical engine indicators are installed on both instrument panels. The oil pressure and manifold pressure indicators read out pressure directly from the engine; the fuel pressure indicator reads out pressure directly from the carburetor. When the engine is not running, manifold pressure reading corresponds to barometric pressure. The tachometer is self-generating and does not require aircraft electrical power input. Oil, cylinder head, and carburetor air temperature indicators, how-ever, depend upon 28-volt d-c power from the primary bus for operation.

North American T-28C Trojan for X-Plane by Khamsin & Arno54


The engine drives a three-blade, constant-speed Hamilton Standard hydrornatic propeller. A double-capacity governor, controlled by mechanical linkage from the cockpit, maintains a selected rpm, regardless of varying airspeeds or flight loads. The governor and oil pump are contained within a constant-speed control assembly mounted on the nose section of the engine. The dome of the three-blade propeller contains a piston-actuated geared cam, which is meshed to the gear teeth on the propeller blades. Engine oil pressure, boosted by the governor pump, moves the piston, which moves the propeller blades to the desired pitch. If the propeller is in an underspeed condition, the governor permits oil to drain from the outboard side of the piston. Then, the centrifugal twisting movement of the propeller blades can move the piston outboard, decreasing the blade angle and increasing the engine rpm. If an overspeed condition exists, high-pressure oil from the governor is directed to the outboard side of the piston. This oil forces the piston inboard, increasing the blade angle and decreasing the rpm. During an on-speed condition, a difference in pressure is maintained by the governor to offset the centrifugal twisting movement of the blades. This holds the blades at the desired pitch. The propeller is governed within the range of 1200 to 2700 rpm. Fast throttle bursts should be avoided with the propeller control set above 2500 rpm. Because of the extremely rapid acceleration of the engine, the engine will overspeed before sufficient high-pressure oil can be supplied by the governor to correct the overspeeding condition.

NOTE Loss of engine oil pressure will cause the propeller blades to go to maximum low pitch (full increase rpm) regardless of propeller control position or the amount of power being generated by the engine. If power is applied at low airspeed with the propeller rpm in full decrease, the engine rpm may decay even if the propeller control is retumed to full increase. Recovery from this undesirable condition may necessitate reducing the throttle and lowering the nose to increase airspeed and rpm to the point that centrifugal twisting moment exceeds aerodynamic loading.


Engine rpm is determined by the setting of the PROP control lever, located on the throttle quadrant in each cockpit. The position of the lever determines the setting of the propeller governor.


Oil for engine lubrication is supplied from a 12.2 U.S. gallon oil tank, Of the total, 8.8 gallons are usable, with 3.4 gallons foam and expansion space, Lubrication is accomplished by a pressure system with a dry sump and scavenger pump return. Oil flows, by gravity, from the tank to the engine pressure pump, which forces it through the engine. Two scavenge pumps force the oil through either t.he oil cooler warmup jacket or through the oil cooler (depending on oil temperature), then back to the tank. The engine cowl and oil cooler flaps are manually controlled and driven by an electric motor. The two cowl Haps are located on each side of the engine cowling and the oil cooler Hap is located on the lower left side of the cowling. When full-open, the oil cooler {lap is open approximately 2% inches more than the cowl Haps. In the closed position, the oil cooler Hap will be approximately 5 degrees open when the cowl flaps are closed. This design balances the temperature dilferential between the cylinder heads and the oil cooler.


Both the cowl Haps and the oil cooler flaps are operated simultaneously by means of a toggle switch on the left console forward of the throttle quadrant in each cockpit. Placing the switch at OPEN or holding it at the spring-loaded CLOSE position operates the cowl and oil cooler flaps. No position indicator is necessary as the cowl flaps are visible from either cockpit. Cowl flaps generate a lot of drag, therefore reducing speed.



North American T-28C Trojan for X-Plane by Khamsin & Arno54



One of the basic limitations placed on engine operation is imposed by the amount of pressure developed in the cylinders during combustion. If this pressure becomes excessive, it can cause detonation and will result in eventual engine failure. Since improper coordination of the use of the throttle and prop lever can cause these limitations to be exceeded, it is important to learn the correct sequence in which these controls should be used, Whenever the engine power is to be reduced, retard lhe throttle first; then retard the PROP lever. Conversely, when increasing engine power, advance the PROP lever first; then advance the throttle.


The MIXTURE control lever on the throttle quadrant is provided with two positions (NORMAL and RICH) for use during flight. The RICH position should be used only for ground operation, take-off, climb, descent, landing, and when operating at normal rated power and above with alternate air; the NORMAL setting should be used during all other conditions of flight. The injection-type carburetor on the R-1820 is equipped with an automatic mixture control to maintain the mixture setting selected, regardless of changes in altitude or temperature. No intermediate position between RICH and NORMAL, or between NORMAL and IDLE CUTOFF, should be selected to arbitrarily adjust the mixture. All performance data in Section XI (except Take Off Distances) are calculated for performance with the mixture control lever in the NORMAL position. Moving the lever between NORMAL. and RICH will increase fuel consumption and decrease planned fuel reserve. Moving the lever between NORMAL and IDLE CUTOFF will lean the mixture and may seriously damage the engine by causing rough engine, backfiring, overheating, detonation, loss in power, or sudden engine failure.


The supercharger low position is used for ground operation, take-off, and during flight up to an altitude where it is more advantageous to operate with high blower (approximately 13,500 feet at military power or 15,000 feet at normal rated power). At the appropriate altitude, before shifting the supercharger to high, reduce the manifold pressure to less than 20 in. Hg and adjust prop control to obtain 1600 rpm. Supercharger shifts from low to high should be made rapidly to avoid wear on the clutch. When a shift is made, be sure the handle is in the HIGH detent position to prevent clutch slippage. A slight increase in manifold pressure will indicate engagement of the high blower clutch. The supercharger should be shifted to low for all descents.



It is often asked what the consequences would be if the 5-minute limit at take-0ff power (rich mixture) or the 50·minute limit at military power (normal mixture) were exceeded. Another frequent inquiry is how long a period must be allowed after the specified time limit has elapsed until take-off power can again be used. These questions are difficult to answer, since the time limit specified does not mean that engine damage will occur if the limits are exceeded. It does mean that total operating time at high power should be kept to a reasonable minimum in the interest of prolonging engine life. In fact, the former procedure may even be preferable, as it eliminates temperature cycles which also aggravate engine wear. Thus, if flight conditions occasionally require exceeding time limits, this should not cause concern so long as constant effort is made to keep the overall time at take-of power to the minimum practicable.Note : The stress of the engine actually depends on the intensity AND the time of exceeding abnormal conditions.


Another factor, to be remembered in operating engines at high power, is that full military power is to be preferred over military power rpm with reduced manifold pressure. This procedure results in less engine wear for two reasons: first, the resulting high brake horsepower decreases the time required to attain the objective of such high-power operation; second, high rpm results in high loads on the reciprocating parts due to inertia forces. As these loads are partially offset by the gas pressure in the cylinders, the high cylinder pressures resulting from use of high manifold pressure will give lower net loads and less engine wear due to piston "slap" as seen at low MAP (manifold pressure). Sustained high rpm is a major factor producing engine wear. High rpm and low manifold pressure require more "rpm minutes" or "piston ring miles" to attain the desired objective of high-power operation.


The engine produces 1425 brake horsepower at sea level for take-off purposes. However, if maximum power is not required, NATOPS procedures recommend using 47 inches MAP with full increase rpm. The "piston miles" necessary to gain a given altitude at this power will normally be less than those obtained at 2700 rpm using a reduced MAP. Take-off power should be maintained until the recommended climbing speed and a safe altitude, considering the local terrain, are attained, but no longer. Cooling the shaft and the engine, afterward, is very recommanded not to risk an engine seizing or shaft explosion.


With high-power operations, it is possible to overtemperature the engine without the overtemperature being observed in the cockpit. FCLP (field carrier landing practice) is an example of this condition; therefore, it is recommended that engine temperature and pressures be closely monitored, particularly during hot weather operations. In particular, cylinder heads are slowly damaged above 210°C whereas 225°C is acceptable. In other words, it's better to be at 225°C for a short period, then to cool to 180°C, than to be at 210°C on the long run.

High-power operations also require high fuel consumption. For this reason, great care must be exercised when operating at high power, especially at low altitude, For example, at military power at 1000 feet altitude, the fuel supply will be exhausted in approximately 45 minutes. Therefore, when practicing take-offs and landing or when flying at high power keep an eye on fuel quantity.



Flight testing has shown that one of the most critical periods, with regard to adequate cooling of the power plant installation, is after the engine is shut down. A large amount of heat is generated during flight, and some of it is retained after shutdown. Since there is no longer any cooling airflow over the engine, the retained heat is conducted throughout the engine. Retained heat may be sufficient to raise the temperature of power plant components above their limits, and serious damage may result. Intake pipes may warp or crack, rocker box covers may warp (resulting in excessive oil leakage), insulation or electrical wires and magnetos or generators may be damaged, etc. For this reason it is very important to idle the engine until cylinder head temperatures drop to at least 150°C, or stabilizes to a value consistent with the existing ambient temperature before shutting down. Cylinder heads will usually be cooled by the time the parking area is reached. This is not always the case, especially in hot weather. The cylinders should always be allowed to cool before the engine is stopped. This explains why an engine can spontanneously catch on fire, in flight, when you have shut the engine off (for training purposes or by necessity.)


  • 1. Refrigeration Ice is usually encountered at temperatures above freezing. It is even possible at OAT above 210°C. The engine is most susceptible to this type ice under conditions of high humidity, and it is posible for it to occur on clear days. The invisible water vapor may precipitate and freeze near the point of fuel injection (discharge nozzle) due to decreased temperatures caused by fuel vaporization. Icing of this type causes a restriction in the induction system manifesting itself as a los of MAP.
  • 2. Impact System Icing generally occurs when the outside air temperature is below 10°C and evidence of visible moisture is present. Any engine instability is fair warning of impending carburetor impact system icing and corrective measures should be immediately taken. Note
    Impact icing may manifest itself as engine instability followed by a loss of power, or where the top deck screen has become iced as gradual loss of MAP. In either event, preventative measures should be taken.
  • 3, Expansion Ice occurs only within a narrow temperature range just above freezing, Moisture in the induction air may precipitate when cooled by adiabatic expansion or passage through carburetor venturi and throttle. Under high humidity conditions near freezing, this condensed moisture frequently freezes as a further result of expansion cooling within the carburetor itself. This condition is most prevalent at near closed throttle positions and manifests itself as a loss in MAP. When conditions are such that throttle icing may be suspected, such as descents during inclement weather at temperatures near freezing or while holding for take-off clearance, preventative pre-heat should be used.

The most important point of preventing a fuel metering disturbance is that of anticipating an icing condition and taking the following action :

  • 1. Set mixture to RICH.
  • 2. Apply carburetor preheat to the minimum value required to prevent icing for any given installation. This will reduce the critical altitude; however, there is no practical substitute under severe conditions. Do not exceed 38°C CAT. Experience indicates 20-30°C is adequate.

If an icing disturbance is encountered without warning the same action should be taken.


The use of carburetor heat is not recommended during take-off. Use maximum allowable heat during ground operation prior to application of take-off power.


In general, icing conditions are usually not encountered above 25,000 feet. At lower altitudes, increasing or decreasing altitude will often effect escape from icing conditions. If icing in any form is encountered and the use of alternate air will not provide sufficiently high carburetor air temperature to remove it, change altitude to get out of the icing level. Use caution in the use of the carburetor air control handle, as extremely high mrburetor air temperatures contribute to detonation and resulting engine damage. ln addition, engine power is reduced by use of alternate air, because airflow is decreased. Full engine power, therefore, is not available with the carburetor air control handle at ALTERNATE. When it is necessary to use altemate air to remove carburetor ice, the carburetor air temperature limits should be closely observed. At all other times, the carburetor air control handle should be in the DIRECT position.

North American T-28C Trojan for X-Plane by Khamsin & Arno54


The aircraft fuel system is entirely automatic after being put into operation. Tuming the fuel shutoff handle ON from either cockpit opens the fuel shutoff valve and starts the d-c powered boost pump in the sump rank located in the right wheel well. Fuel flows by gravity from the wing tanks into the sump tank automatically maintaining an equal level in each wing unless either (a) the fuel tank vents under the flaps are creating unequal pressure or (b) the aircraft is flown out of balanced flight for an extended period. Uneven fuel flow can possibly be corrected in flight by slipping the aircraft so that the balance ball is pushed out toward the wing with the lower fuel level. The boost pump forces fuel under a pressure of 19 to 24 psi through the fuel shutoff valve, the strainer, and the engine-driven fuel pump. The engine-driven fuel pump then boosts the fuel to an operating pressure of 21 to 25 psi. lf the engine—driven fuel pump fails, the boost pump will still supply sufficient fuel for satisfactory operation. If failure of the boost pump occurs, the enginedriven fuel pump will enable the engine to operate up to 10,000 feet.

Two interconnected fuel cells are located in each wing. Port and starboard overboard vents are located in each wing flap to assist even fuel flow and to vent the tanks. As with most fuel systems of this type, prolonged unsymmetrical flght may cause uneven fuel flow or prevent fuel from Howing into the sump tank. If this occurs and the engine stops due to lack of fuel, immediately revert to normal flight and the engine should start.



A fuel shutoff control handle, located on the left console of each cockpit, has two positions: ON and OFF. Each position operates the fuel shutoff valve and the boost pump simultaneously. No action by the pilot is necessary to maintain an equal fuel level in each wing other than normal flight and co-ordinated turns.


A fuel boost pump test switch is located on the electrical switch panel in the forward cockpit. The test switch is wired in series with the boost pump switch on the fuel shutoff control handle. When held in the TEST position, power to the boost pump is interrupted, allowing the engine-driven fuel pump pressure to be checked.


A fuel quantity indicator is located on the main instrument panel in each cockpit. The indicator system is electrically operated and measures the total fuel supply in pounds. The system automatically compensates for changes in fuel density so that the quantity indicator will always register the actual number of pounds of fuel in the tanks, regardless of fuel expansion or contraction due to temperature variation. The full mark on the indicator is set at 1040 pounds for a 25°C (77°F) fuel temperature. If the tanks are full but the temperature is above 25°C (77°F), the indicator will indicate less than full; if the tanks are full and the fuel temperature is below 25°C (77°F), the indicator will indicate above the full mark.

Note Fuel range per pound is the same regardless of fuel temperature.


A low-level warning light is located on the instrument panel in each cockpit. If fuel quantity falls below approximately 100 pounds, the light illuminates, lf the fuel quantity indicator is not operating, the fuel low-level warning light should not be interpreted to mean there are exactly 100 pounds of fuel available, since its indication is only approximate.



The 28-volt d-c power system is powered by a 28-volt, 200-ampere engine-driven generator (30-volt, 300 ampere*) and a 24-volt storage battery serves as standby power. D-C power can also be supplied through an external power receptacle.


Power for the a-c system is supplied by two inverters: a 750-volt-ampere inverter and a 250-volt-ampere inverter, both being operated at once through the switch on the right panel.


D-C power is distributed from 2 buses: battery + primary, and secondary (external). Both buses are powered from the generator through the d-c power switch. The primary bus is also energized by external power, and by battery power when the DC Powiax switch is in either BAT. , GEN. or BAT. ONLY position. The secondary and monitored buses are energized through the primary bus by generator or external power. The secondary bus is energized by the battery + generator when the "ext" switch is "on". The secondary bus drives the exterior power of the plane (pitot and bulbs), and the primary all the rest.



The DC POWER switch, located on the right forward console in each cockpit, has BAT. & GEN., OFF, and BAT. ONLY positions. The switch cannot be moved to the BAT. ONLY position when the guard is down. With the switch at BAT. or GEN. position and the generator operating, power is supplied to all four d-c buses. With DC POWER switch off and the generator operating, the battery bus is energized by the battery and all other buses are inoperative. With the DC POWER switch at BAT. at GEN. and the generator not operating, the battery and primary buses are energized by the battery, and if the landing gear is extended, the secondary bus is automatically connected to the system. With the switch at the BAT. ONLY position and the generator inoperative, the battery, primary, and secondary buses are energized by the battery. Only the battery bus is energized if the switch is orr and the generator is not operating.


  • The DC POWER switch should not be placed at BAT. & GEN. or BAT. ONLY when external power is applied.
  • If it is necessary to use the battery simultaneously with external power, battery "on" time should be kept to a minimum since no ground ventilation is provided.

The front cockpit DC POWER switch is marked: BAT. & GEN., OFF, and BAT. ONLY. The rear cockpit switch is a lever-lock type with two positions marked: NORMAL. ON and EMER. OFF. The front cockpit switch cannot be moved to the BAT. ONLY position unless the guard is raised. The rear switch can be moved to either position by pulling up on the toggle before moving. With the front switch at BAT. & GEN. position and the generator operating, power is supplied to all four d·c buses. With this switdm ore and the generator operating, the battery bus is energized by the battery and all other buses are inoperative. With the DC POWER switch at BAT. & GEN. position and the generator not operating, the battery and primary buses are energized by the battery and, if the landing gear is extended, the secondary bus is automatically connected to the system. With the switch in the BAT. ONLY position and the generator inoperative, the battery, primary, and secondary buses are energized by the battery. With the switch on and the generator not operating, only the battery bus will be energized. For normal operation, the d-c switch in the rear cockpit should always be in the NORMAL ON position. If an emergency occurs, placing the switch in the EMER. OFF position will cut of all d-c power in the system except the buttery bus.


The two-position instrument power switch, located on the right console in each cockpit, controls the bus load from the main and the standby inverters. The switch can be operated only by the pilot who last actuated his control shift switch. The main inverter supplies current to the a-c bus when the switch is at the No. 1 INV. position, and the standby inverter operates simultaneously under a dummy load. Moving the switch to N0. 2 INV. position connects the standby inverter to the a-c bus, and the main inverter then operates in an open circuit.


The instrument power switches have been removed from the transfer system. The aft cockpit switch cannot be moved to the N0. 2 INV. position unless the guard is lifted. The front cockpit inverter selector switch is used to select the power source, main or standby inverter, to all a-c equipment except the aft cockpit attitude gyro. The aft cockpit inverter selector switch is used to select the power source to the aft cockpit attitude gyro and to the rear cockpit. instrument power failure relay. If TACAN is installed, it is powered by the 750-volt-ampere standby inverter only, and controlled by a relay which is actuated when both inverter selector switches are in the No. 1 position. If either switch is moved from the No. 1 position, power to the TACAN will be interrupted and the equipment will be inoperative.


An intercockpit control shift switch is mounted on both right consoles . The control shift system is energized by the battery bus and will function with the DC Power: switch ore in both cockpits. lf the DC rowcx switch is positioned to either snr. at GEN. or mrc. outy, or if external power is supplied, a light adjacent to the control switch in either cockpit marked tr. ON cormzot will illuminate when the related switch is operated to take control.


All d-c circuits are protected from overloads by push-to- reset circuit breakers. Should an overload occur in a circuit, the resulting heat rise causes the circuit breaker to pop out and open the circuit. The circuit breaker may be pushed in again in an attempt to re-energize the circuit. However, the circuit breaker should not be held in if it opens the circuit a second time. The circuit-breaker panel is located in the front cockpit below the right console. Alternating-current circuits are protected by fuses. The fuse panel can be reached through an access doot in the front cockpit above the right console.


A generator-off warning light is located on the right forward console in both cockpits. Illumination indicates that the generator is not operating and the battery or external power unit is supplying all power for the electrical system. To conserve the battery when the generator is inoperative, all unnecessary electrical equipment should be turned off or disconnected by pulling out related circuit breakers.


A flight instrument power failure warning light, mounted on both instrument panels, illuminates when the a-c bus is not energized or power is interrupted to the attitude gyro. Illumination with the instrument power switch at No. i INV. position indicates that the attitude gyro or the main inverter is not operating. If the main in- verter is inoperative, the attitude gyro and other a·c powered equipment will be inoperative. lf the light goes out when the instrument power switch is moved to the N0. 2 INV. position, a—c power is being supplied to the bus by the stand- by inverter and failure of the main inverter is confirmed. Failure of the stand-by inverter also causes the light to illuminate, providing the switch is in the N0. 2 INV. position. lf the light remains illuminated, regardless of the position of the instrument power switch, either both inverters are inoperative or the attitude gyro is not receiving a-c power, lf one of the ATT GYRO & INST FAIL IND fuses is blown, the attitude gyro will be inoperative, but all other a-c powered equipment will be operative even though the waming light remains illuminated. Aircraft with ASC No. 36 incorporated have separate fuses for each cock- pit attitude gyro and separate warning light circuits.


A voltmeter, located on the instrument panel, indicates voltage output of the generator battery, or external power unit. Normal voltage indication is approximately 27.7 volts.


A generator load indicator is mounted on both instrument panels. The indicator reflects the percent of generator output being used and is graduated in decimal fractions. An indication of 0.5 means the electrical system is using one-half of rated generator capacity.


Hydraulic power is used to operate the landing gear, wing flaps, canopy, speed brakes, and, in T-28C aircraft, to retract the arresting hook. A variable-displacement, engine-driven pump supplies hydraulic pressure. When no hydraulic units are being operated in flight, the entire output of the pump is automatically diverted to the hydraulic reservoir through an electrically actuated bypass valve. When any hydraulic control is operated, the bypass valve is electrically deenergized and closes, allowing the hydraulic system to build up pressure for operation of the selected unit. The bypass valve is electrically energized to open and depressurizes the system only when all units are in their normal flight position, and the system is pressurized at all other times. System pressure is maintained whenever the speed brake or canopy are open and whenever the gear, T-28C arrest- ing hook, or flaps are in any position other than up and locked. In the event of electrical failure, the bypass valve automatically closes, pressurizing the system. A standpipe in the reservoir retains enough fluid to supply the wheel brakes if all fluid is lost from the reservoir. A high- pressure air system is provided as an altemate for the hydraulic system to open the canopy in an emergency. A hydraulic pressure gage (8, figure 1-3) is located on the left console in each cockpit. The hydraulic system is d-c controlled from the secondary bus. For hydraulic system servicing, refer to Section I, Part 3.


A hydraulic handpump is provided in the front cockpit. The handpump is used primarily for ground check of the hydraulic system, but may be used in flight should the engine-driven pump fail. The hand-pump is merely a substitute for the engine-driven pump and does not provide separate fluid supply or lines to operate any part of the system.


The primary flight control surfaces (elevator, ailerons, and rudder) are operable from either cockpit by interconnected stick and rudder pedals. No hydraulic boost control system is provided. Cable-operated trim tabs on all control surfaces, except the right aileron tab, are manually positioned from trim tab control wheels located on the left consoles. Hydraulically operated, semislotted wing flaps are mounted on the trailing edge of each wing and are controlled by a lever on the throttle quadrants. A single, hydraulically operated, perforated speed brake, mounted on the bottom of the fuselage, may be extended at any speed. The rudder pedals are adjustable fore and aft and incorporate wheel brake control by pressure on the top of the pedals. All primary flight controls can be locked in a neutral position by a mechanical control lock in the front cockpit. This lock also secures the throttle in the closed position.


The control stick in both cockpits incorporates positive grip handles. The stick grip in the front cockpit contains a gun trigger, a bomb release button marked B on top of the grip, and a rocket firing button marked R on the side of the grip. The control sticks are mechanically connected to each other by a push-pull tube. The elevator system incorporates a bungee and bobweight. The bungee provides satisfactory control "feel" during low-speed flight and landing, and the bobweight assists pilot effort during accelerated flight.


A set of interconnected rudder pedals in each cockpit controls the rudder action through direct mechanical linkage to the rudder. The wheel brakes are actuated by pressure on the top rudder pedal plates. The rudder pedals can be adjusted for correct leg length.

Rudder Pedal Release Levers

A rudder pedal release lever, located below the center of the instrument panel in each cockpit, permits individual adjustment of the pedals, Holding the lever to the right unlocks the pedals, allowing them to move full aft. Releasing the lever locks the pedals in the desired position. A cable running between the pedals prevents unequal adjustment.

As the rudder pedals are spring-loaded to the aft position, have feet on the pedals when making adjustments.


Cable-operated elevator, rudder, and aileron trim wheels are located on the left console in each cockpit. Trim tab position is shown by a scale and pointer at each control. The left aileron trim tab is controlled by the aileron trim tab wheel. The right aileron trim tab can only be adjusted on the ground.


All surface controls and the throttle are locked by a control lock, which is stowed on the door of the front cockpit behind the stick. Pulling the plunger- type pin on the` right side of the lock releases it from the stowed position on the door. The lock can then be raised to engage the stick and rudder pedals in the neutral position. The plunger is then released to engage a hole in the side of the stick. VU hen subsequently closed, the throttle is locked and will remain locked until the control lock is released.


Hydraulically operated, semislotted wing flaps extend from aileron to fuselage on each wing. The flaps are operable from either cockpit and a flap position indicator is provided on each instrument panel. No emergency system is provided for operating the Haps. However, if the hydraulic system fails, an attempt may be made to extend the flaps by operating the hydraulic handpump. The handpump will pressurize the system only when loss of pressure is caused by failure of the engine—driven pump. The Haps can be lowered manually on the ground from outside the aircraft and spring-loaded doors in the surface of the flaps may be used as steps up to the wing.



The daps are operated by means of a FLAP handle on the throttle quadrant in each cockpit. The handle is shaped in the form of an airfoil for easy recognition by feel. Moving the handle to UP, 1/4, 1/2, 3/4, or DOWN electrically closes the solenoid bypass valve, which pressurizes the hydraulic system to operate the flaps, and mechanically positions the wing flap selector valve. When the desired flap position is obtained, the selector valve automatically returns to the neutral position. The flaps will lower 37% degrees when the FLAP handle is moved to DOWN. Detents hold the handle in the selected position. The flaps require approximately 7 to 12 seconds to extend and approximately 10 to 15 seconds to retract. The hydraulic system is pressurized as long as the flaps are in any position other than up and locked. The flaps are held in the up position by a mechanical overcenter lock and hydraulic lock and in any selected down position by the hydraulic locking action of the selector valve in the neutral position.


The manual flap lever is located on the left side of the fuselage, above the wing trailing edge, The lever is provided to release the flap uplock and to open a bypass valve on the actuator, allowing the flaps to be pushed down manually to 50 degrees, enabling the pilot to reach the steps. If the engine is running when the lever is pulled, the flaps will lower hydraulically to 37% degrees, but must be manually pushed down to 50 degrees (with the lever still held out). When the lever is released, the flaps will return to 37% degrees if hydraulic pressure is available; the steps should not be used unless the flap is fully extended to 50 degrees. Lowering the flaps from the outside also moves the FLAP handles in the cockpits.


A flap position indicator , calibrated in degrees, is located on the instrument panel in each cockpit. Normal hydraulic Hap travel is 37 1/2 degrees to the full down position. The flaps can be lowered manually to 50 degrees on the ground for access to the steps in the flaps. The flap position indicators are d-c powered from the secondary bus.


The hydraulically operated speed brake is essentially an additional flight control which is useful for making descents or moderate deceleration from high speeds. The brake can be opened at any airspeed up to maximum and, although brake opening causes a nose-up pitch, the forward stick pressure necessary to maintain the desired aircraft attitude is moderate up to 250 knots IAS. Above 250 knots, speed brake extension causes a nose-up pitch which requires large initial stick pressures to maintain a constant dive attitude. This stick pressure can be trimmed out at all speeds by adjustment of the elevator tab. Because of this excessive nose-up pitch, speed brakes should not be extended at the initiation of or during a high·speed pullout. Speed brakes should be extended prior to entering high-speed dives. Failure to observe these precautions may result in overstress during a pullout. Closing the brake, of course, causes a nose-down pitch. A limiter valve is incorporated in the speed brake hydraulic system to reduce the violence of the nose-up pitch and preclude possible overstress when the speed brake is opened at high airspeeds. This valve reduces the pressure of the hydraulic fluid to the speed brake actuating cylinders and automatically allows the speed brake to extend to variable openings, dependent upon airspeed. Variable opening positions are maintained until the airspeed is reduced sufficiently to allow the limiter valve to overcome the force created by aerodynamic pressure and to fully open the speed brake. Conversely, if the speed brake is fully extended and high airspeed is attained, the limiter valve will automatically adjust degree of opening until airspeed has been reduced. The speed brake will then fully extend. When the speed brake is in use, the pilot has not positive control over the limiter valve other than airspeed control. Stick pressure to maintain a constant dive angle can be trimmed out at all speeds by adjustment of the elevator tab. When closing the speed brake, the limiter valve has no function in the system and the normal nose-down pitch will be experienced.


A speed brake switch is located on top of the throttle grip in each cockpit, The speed brake can be operated by the pilot who last actuated his control shift switch. The speed brake switch has only two fixed positions, OFF and ON, and no intermediate positions can be selected. When the switch is moved to ON, it energizes the speed brake selector valve, electrically closes the solenoid bypass valve, and pressurizes the hydraulic system. The hydraulic system remains pressurized as long as the brake is open. In event of an electrical failure while the speed brake is open, the speed brake automat- ically closes to a trail position, depending on airspeed. Should hydraulic failure occur while the brake is open, the brake will stay open until the speed brake switch is moved to oss; air loads will then close the brake to a trail position.

CAUTION Do not operate speed brake while the baggage compartment door is open. With the baggage compartment door open, the aft edge of the speed brake will not clear the baggage compartment door.


The retractable tricycle landing gear is hydraulically operated. The main gear retracts inboard into the wing and fuselage: the nose gear retracts aft into the fuselage. Mechanically operated fairing doors cover the wheels in the retracted position. All fairing doors remain open when the gear is down. The gear is held up by mechanical locks and held down by overcenter side brace lockpins on each gear and by hydraulic pressure. The lockpins will hold the gear down should the hydraulic system fail. All uplocks are released by initial movement of the landing gear handle; consequently, in event of hydraulic failure, the gear can be unlocked by the gear handle. Once released, the main gear extends by its own weight, and the nose gear is extended fully by a spring bungee. A solenoid-operated downlock in the landing gear control system prevents inadvertent gear retraction when the aircraft is on the ground. Any landing gear position other than up and locked deenergizes the solenoid bypass valve and causes the hydraulic system to be pressurized. The full-swiveling nose wheel is equipped with a shimmy damper and automatic airborne centering, Brakes on the main wheels are used for directional control until the rudder becomes effective. A fixed tail skid is installed under the aft section of the fuselage.



The landing gear handles are located at the left of the instrument panel in each cockpit. Moving either handle to UP or DOWN operates the gear selector valve, deenergizes the bypass valve which pressurizes the hydraulic system, and mechanically positions the gear uplocks. When the aircraft is on the ground, a solenoid-operated downlock (deenergized) prevents normal movement of the gear handles out of the DOWN position. The lock is automatically disengaged (electrically energized) when the aircraft is airborne. If an emergency requires gear retraction on the ground, pulling either handle up very sharply to UP overrides the solenoid-operated lock. In the event of hydraulic failure in flight, the gear can be lowered by moving the gear handle to DOWN, which mechanically releases the uplocks, allowing the main gear to drop. It may be necessary to yaw the aircraft to lock the main gear down. The nose gear is forced down and locked against air load by a spring bungee. With normal hydraulic system operation, the landing gear extends in approximately 4 to 6 seconds and retracts in approximately 6 to 10 seconds. To lock the gear handle at FULL UP or DOWN, it must be placed firmly in its extreme position. Any time the landing gear is not locked in the position required by the gear handle, a red light in the gear handle illuminates. Refer to LANDING GEAR AND WHEELS WARNING LIGHTS, in this section. Because of the gear handle location in the down position, care must be taken to avoid a gear retraction by inadvertently striking the handle with the knee.

Note With the engine windmilling, prop control lever at full DECREASE RPM, the hydraulic pump will maintain 1500 PSI only until a system is actuated. Once actuated a decay in pressure occurs waich in the case of the landing gear requires a slightly longer retraction period (7 sec.). Landing gear free fall results in a shorter extension time (4 secs.).


Position of the landing gear is shown by three individual indicators, one for each gear, located on the instrument panel in each cockpit. Each indicator shows crosshatching if the related gear is in any unlocked condition and in the absence of electrical power, The word UP appears if the gear is up and locked, and a wheel shows if the gear is down and locked. The indicators are d-c powered from the secondary bus.

Note The landing gear position indicators may stick and not fully return to the crosshatched deenergized position when the electrical system is turned off. However, the indicators will return to normal operation immediately when the electrical system is energized.


Landing gear ground safety Iockpins are provided for insertion in the side brace of each of the three landing gear struts to prevent accidental retraction of the landing gear on the ground. The pins must be removed before flight or the gear will not retract.


Additional warning of unsafe gear position is provided by a red light incorporated in the landing gear control handle and a flashing WHEELS warning light located at the top left corner of each instrument panel (5, figure 1-4). The WHEELS light will illuminate and flash whenever the throttle is retarded below approximate minimum cruise setting, with the flaps in any position other than fully retracted, and the landing gear not locked down. The WHEELS light will also provide a warning indication with the flaps fully retracted if the propeller control lever is at or above the 2500 rpm position. On aircraft equipped with the gear warning horn, the hom will be actuated when the throttle is retarded to approximately 15 inches MAP, regardless of flap position, if the landing gear is not down and locked. A horn silencer button is available and is located at the base of the throttle quadrant (figure 1-3). Both warning systems are d·c powered through the primary bus.

Caution Some aircraft do not have the 2500 rpm propeller control lever switch installed. ln these aircraft, when practicing no flap landings, the warning lights will not flash (flaps full up) and the only unsafe landing gear indication will be the position indicator and the waming light in the landing gear handle.


To aid in determining landing gear position from the ground at night, a white light is installed on each gear strut. Each light illuminates only when the related gear is down and locked and the EXT MASTER lights switch is turned 0N.


An arresting hook is installed on T-28C aircraft only. The hook is extended by bungee pressure and gravity and is retracted by hydraulic pressure. The hook can be lowered under most emergency conditions, as the uplock is mechanically connected to the control handle and the bungee action is positive and independent of any other system. If a hydraulic failure occurs and it is necessary to retract the hook, this can be attempted by placing the control lever in the UP position and operating the handpump. If there is hydraulic fluid in the reservoir and the hook retract system is intact, the hook should retract. An electrical failure will not alfect hook operation in any way except for the unsafe indicator lights.


An arresting hook handle is located to the right side of each instrument panel, and is interconnected between cockpits. The handle has two positions marked UP and DOWN, and movement of the handle is in a vertical plane.


The arresting hook unsafe light, incorporated in the hook handles, receives power from the primary d-c bus. Only a complete electrical failure renders the light inoperative. The lights will illuminate whenever the hook is in any position other than full down or locked up. When the hook is moving from one position to the other, the light will be on.

Caution If the light does not go our when hook travel is completed, do not attempt an arrested landing until hook is visually checked for a safe condition.


The main wheel brakes are of the manual, hydraulic, master cylinder type, operated by toe pressure on the rudder pedals. No boost is supplied by the aircraft hydraulic system, but fluid from the hydraulic system reservoir supplies the master cylinders. Should all fluid be lost from the reservoir, adequate fluid remains in the standpipe and lines to supply the brakes for normal operation. No emergency method of operating the brakes is provided. A parking brake handle (24, figure 1-4) is installed in the front cockpit only, to the right of the instrument panel. Parking brakes are set by depressing the pedals, rotating the parking brake handle, and then releasing pedal pressure. Brakes are subsequently released by depressing the pedals (front cockpit only).

Caution Possibility of complete brake failure exists following heavy braking, due to vaporization of fluid in the overheated brake assembly.


The aircraft displays very little aerodynaruic buffet prior to stall in the power approach and waveoff configurations; therefore, a mechanical stall warning device is incorporated to warn of impending stall approximately 7 or 8 knots above actual stall speed. This system consists of a transducer mounted on the right wing leading edge, a lift computer located in the nose wheel well, a rudder pedal shaker on the right rudder pedal in each cockpit, a landing gear switch on the right main gear strut, and a test switch on the left forward console in the front cockpit. The stall warning system functions on the principle that as aircraft angle of attack changes, so does the pressure distribution on the wing. The transducer is located so that just prior to the stall it senses the change of pressure and actuates the circuit to the rudder pedal shakers. The lift computer balances the circuit to maintain the margin of stall warning within the range between 105 and 115 percent of stall speed. Whenever the weight of the aircraft is on the landing gear, the switch on the right gear strut disrupts the circuit to make the pedal shakers inoperative. Operating the test switch on the left forward console will bypass the ground safety circuit and the pedal shakers will operate for a preilight check. The transducer is electrically heated when the pitot heater switch is 0N.

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