1. Engine Systems J-52-P408 &J-52-P6/P8

2. To familiarize the Transition Pilot with basic operation to include:
Lesson Objectives
Introduce the major components of the J-52-P408 and the J-52-P8 TURBOJET ENGINE in the A-4N and TA-4.
To familiarize the Transition Pilot with basic operation to include:
Basic Engine Operation
Associated ENGINE instruments & COCKPIT CONTROLS
Operational procedures
Engine Limits
Emergency Procedures
OBJECTIVE:
To familiarize the Transition Pilot with the P&W J52-P6 and the J-52-P408 TURBOJET ENGINEs, basic differences, associated ENGINE instruments, ENGINE operating limitations, and actions necessary in event of ENGINE malfunctions for both engines.
ENABLING OBJECTIVES:
1 State the major sections of the Basic J52 ENGINE.
2 Describe the Basic J52 Engine AXIAL FLOW COMPRESSOR to TURBINE configuration.
3 State the post-starting warm-up and pre-shutdown cooling periods for the J52-P6/P408 ENGINE.
4 Explain the J52-P6 and J-52-P408 IGNITION SYSTEM.
5 State forward cockpit location of the ENGINE instruments:
a EXHAUST GAS TEMPERATURE INDICATOR
b. TACHOMETER
c. PRESSURE RATIO INDICATOR
d. OIL PRESSURE INDICATOR
e. OIL QUANTITY INDICATOR SWITCH.
6 State the sources for indications on the ENGINE instruments.
7 State the power sources for operation of the ENGINE instruments.
8 State the J52-P6 RPM equal to 100 percent. State the J52-P408 RPM equal to 100 percent
9 Recognize the significance of RPM, FUEL FLOWMETER EGT, or EPR fluctuations.
10 Explain the ENGINE OIL SYSTEM.
11 State the starting, idle, and normal operating OIL PRESSURE limits for the A/TA-4 aircraft.
12 Explain operation of the OIL QUANTITY INDICATOR SWITCH.
13 State required pilot action in event of inflight OIL PRESSURE malfunctions.
14 State when EGT should be used as a basis for setting thrust
15 State what determines allowable time limit at a thrust level
16 State the maximum allowable speed in percent and RPM for the J52-P6 ant the J52-P408 ENGINE.

3. Major Assemblies Major Assemblies
The P&W J52-P408 and P6 engines are axial–flow, dual compressor, gas turbine engines with through–flow combustion chambers arranged in an annular chamber, and a two-stage reaction turbine. The five–stage, low pressure compressor is connected by a through-shaft to the second-stage turbine.
The seven–stage, high pressure compressor is connected independently by a hollow shaft to the first stage turbine.
The rpm of the high pressure rotor (N2) is controlled by the engine fuel control, whereas the rpm of the low pressure rotor (N1) is mechanically independent and is a function of the pressure drop across each of the two turbine stages. The P408 engine has a two position, low pressure compressor inlet guide vane assembly, while the P6 inlet guide vanes are fixed. The first stage turbine inlet guide vanes and rotor blades are cooled with high pressure bleed air. There are nine combustion chambers (No. 1 at the top) which incorporate features for reduced exhaust smoke. Spark igniters are located in the No. 4 and 7 combustion chambers.
The J52–P408 engine is rated at 11,200 pounds thrust. This is the thrust the engine will develop during sea level static operation of a standard atmospheric day, with a bellmouth inlet duct and no aircraft accessories installed. The J52–P6 engine is rated at 8,500 pounds thrust.
Also included as part of the engine is the compressor inlet anti–icing system, intercompressor stall air bleed system, selfcontained lubrication system, fuel system, ignition system and fuel heater.

4. Inlet Guide Vane (A-4N Only)
Inlet Guide Vanes
The engine incorporates a compressor inlet stator control system which is designed to assist the engine in adapting to a wide range of operating conditions. The system consists of an inlet stator ratio control, venturi, control valve, two vane actuating cylinders, linkage, and 14 two-position inlet vanes. The movable portions of the inlet vanes compose the rear half of the vanes. The vane movement is synchronized through a series of actuating arms attached by pins to a synchronizing ring. The vanes are actuated by two fuel-pressure-operated cylinders at the upper left and lower right quadrants of the inlet case. Adjustable actuating cylinder stops determine the extremes of vane travel. Pump discharge fuel pressure is directed to the cylinders by the control valve, the positioning of which is determined by a signal from the inlet stator ratio control.
The control valve is a pneumatically operated hydraulic servo valve. In its free position, when vented to a signal pressure, the valve directs fuel pump discharge pressure to the vanes closed side of the vane actuating cylinders and vents the vanes open side to the fuel pump inlet. When the valve receives a signal pressure, it is positioned by a diaphragm to direct fuel pump discharge pressure to the vanes open side of the actuating cylinders, and the vanes closed side is vented to the fuel pump inlet.

5. Engine Anti-Icing System
TA-4
A-4
Right Outboard Wedge Panel
ANTI-ICE Switch
The Engine Anti-Icing System
The engine anti-icing system is designed to prevent ice formation. Safe operation requires that the pilot anticipate the possibility of ice formation whenever these weather conditions exist.
Ice formation is prevented in the engine air inlet section by an integral powerplant system that utilizes hot high pressure bleed air from the compressor section. Bleed air from the top left side of the compressor discharge is piped forward through external lines and distributed through the hollow inlet guide vanes and nose cone from which it is ported into the engine inlet airstream.
Note
Operation below approximately 75 percent rpm may not supply sufficient heat to keep the engine air inlet ducts clear of ice.
Anti-Icing Control
Electrical control of the anti-icing system is accomplished by placing the anti-icing switch, located outboard of the right console, in ALL position. When the switch is in ALL position, power is directed to the electrically positioned (open close) anti-icing control valve and regulator mechanism on the external lines. The heating element of the pitot tube is also actuated in the ALL position.

6. J-52-P408/P6 COMPARISON
Note: Minimum RPM airborne for P-8 is 70%

7. Engine Lubrication System
The engine lubrication system is a self-contained, high pressure system that supplies lubrication to the main engine bearings and to the accessory drives. Oil delivered by the engine-driven oil pump is cooled by means of an oil cooler prior to entering the bearing compartments. The oil cooler is a heat exchanger, employing the fuel flowing to the engine as a coolant. A scavenger system returns oil withdrawn from the bearing compartments and the accessory drive gearbox to the oil tank. A breather system connects the individual bearing compartments and oil tank with the breather pressure relief valve. The breather pressure relief valve vents overboard on the starboard side of the aft fuselage. The maximum oil consumption is .28 gallons per hour (approximately 1 liter per hour).

8. OIL PRESSURE
Oil Pressure Indicator: An engine oil pressure indicator is located on the instrument panel. Normal oil pressure is 40 to 50 psi. Minimum oil pressure for ground IDLE is 35 PSI.
Note
· Maneuvers producing acceleration near zero “g” may cause a temporary loss of oil pressure. Absence of oil pressure for a maximum of 10 seconds is permissible.
· Oil pressure indications are available on emergency generator.
The oil transmitter measures engine oil pressure and converts the information to an electrical output signal. Oil pressure is sensed by a diaphragm separating two pressure chambers. The first pressure chamber measures engine breather pressure while the second chamber pressure varies its pressure in accordance with engine oil pressure. The output is not representative of total oil pressure through the engine lubrication system, but of the difference between gearcase pressure and system pressure. This method provides a fixed value for monitoring engine lubrication system pressure at any altitude or barometric pressure.
In the A-4N an OIL PRESS warning light is located on the left annunciator panel. This light will illuminate when oil pressure drops below operating conditions. (35psi)
Illumination of the this light will simultaneously illuminate the MASTER CAUTION light.

9. Oil Level Warning System. TA-4 Only
OIL QUANTITY
Oil Level Warning System. TA-4 Only
The engine is equipped with an oil tank having a useable capacity of 3.4 gallons. The oil tank is filled by a pressurization filling system. Oil is pumped into the tank until it reaches the fill line on the top of the oil tank and out through the overflow port in the tank to the drain adapter.
The system requires both 115-VAC and DC Primary bus power supplied from either the main aircraft or the emergency generator. The sensing elements (sensors) are thermocouple devices that generate voltage when heated. Under normal conditions, the sensors are submerged in oil and the heat is dissipated within the oil. With electrical power supplied to detector, the unit is automatically connected to the lower sensor in the oil tank. If the oil level is below the lower sensor, heat increases at the sensor and sufficient voltage to switch detector is transmitted from the sensor to the detector, completing circuit, and the OIL LOW light comes on.
The circuit works in a similar manner when the high level sensor is not submerged in oil. The OIL LOW lights will not come on unless the relay is energized by pressing the OIL LOW light. With relay energized and the high level sensor out of the oil, the OIL LOW lights will come on, indicating that less than 80 percent of oil remains in the oil tank.
The Press-to-Test (PTT) will verify the 20% sensor, while the 80% sensor is tested by the PTT on the OIL LOW and MASTER TEST at the same time.

10. Oil Level Warning System. A-4N only
The Oil reservoir system is the same as in the TA-4 aircraft. The warning system consists of warning lights vice the oil quantity cube.
Two separate oil quantity measurement systems in the oil tank are associated with two oil quantity lights on the left annunciator panel.
When oil quantity falls below the 80% level the OIL LEVEL HIGH light will illuminate.
When oil quantity falls below the 20% level the OIL LEVEL LOW light will illuminate.
Both lights will activate the MASTER CAUTION light simultaneously.
Similar to the TA-4 the OIL LEVEL HIGH light may erroneously indicate low oil quantity momentarily during periods of takeoff acceleration.
On start a HIGH OIL LEVEL light is not abnormal and is only indicative of actual oil level after the oil system has been scavenged.
Oil Press Light
INOP

11. A-4 ENGINE IGNITION SYSTEM
The ignition timer energizes the igniters for a 30 to 45 second firing cycle.
The momentary ignition switch, which is used to energize the ignition timer, is the switch actuated by movement of the throttle outboard.
The engine ignition system consists of two spark-igniters, an ignition timer, an ignition switch, and two ignition exciter units. The spark-igniters are located in the No. 4 and 7 combustion chambers, and each is energized by a separate exciter. The ignition timer energizes the igniters for a 30 to 45 second firing cycle. The exciters fire both spark igniters for ignition to start the engine. The ignition switch that is used to energize the ignition timer is a momentary control switch actuated by movement of the throttle. There is no continuous ignition provided on either aircraft.
The ignition system is powered from the Primary AC and DC buses in both the A-4 and the TA-4 aircraft. The Primary buses are normally powered from the main generator or the Emergency Generator during airborne airstart attempts. Both the J-52-P6 and P408 require an external source for both AC and DC power on deck.

12. ENGINE FUEL SYSTEM
-The engine fuel system is designed to permit selection of desired engine power requirements at all operating altitudes and temperatures.
-The system has two major components: The engine fuel pump and the fuel control.
-Fuel is supplied to combustion chambers at pressures and flows necessary to meet these requirements.
-Fuel is supplied from the tanks through necessary strainers and valves to engine driven fuel pump.
-Fuel then passes through a fuel anti-icing system installed between the boost and main stages of the pump and is pumped to fuel control unit where it is metered in proper quantities.
-Excess fuel is returned to the pump.
Engine Fuel System
The function of the engine fuel system is to supply and regulate the fuel to the combustion chambers at pressures and flows required by engine air flow at all operating altitudes and temperatures. The system has two major components: the engine fuel pump and the fuel control.
Engine Fuel Pump
The engine fuel pump consists of a centrifugal booster stage and a high-pressure single gear stage with a 40-micron filter between the two. The boost impeller, in addition to increasing pressure, serves to deaerate fuel and eliminate pump cavitations in the gear stage. The boost impeller is driven at approximately 2 1/2 times the speed of the gear stage.
A filter bypass, a pressure relief valve, and a vapor return to the fuselage fuel cell are also incorporated.
1. A relief valve set at approximately 90 psi above maximum discharge pressure.
2. Fuel bypassed from the relief valve, filter bypass, actuating cylinder control valve, or from fuel control returns to inlet side of gear stage.

13. FUEL ANTI-ICING
-When fuel in the aircraft fuel tanks becomes severely chilled, icing of the filter may occur, and obstruct fuel flow. A bypass valve will allow fuel to bypass filter if filter is obstructed.
-Fuel anti-icing system prevents this icing condition by automatically warming fuel, using hot compressor discharge air.
-Although fuel flows through heater at all times, airflow is prevented unless conditions require fuel heating.
-Fuel temperature sensing element of air shutoff valve senses and controls fuel temperature at approximately 50 deg F (10 deg C)
The fuel anti-icing system basically consists of fuel heater and core assembly, anti-icing air regulator valve, fuel heater bypass valve, air shutoff valve and the air outlet valve.
Fuel discharges from the centrifugal-boost stage of the engine fuel pump and flows around heater air tube passages, to the temperature sensing elements of air shutoff valve and air regulator valve, and then to filter in the fuel pump.
A heater fuel bypass relief valve is provided to bypass fuel if the heater fuel passages become restricted.
Hot Ps4 air flows directly from the engine diffuser case past the air shutoff valve and through heater air tubes, and is bled overboard.
The air shutoff valve is normally closed, opening only when fuel temperature decreases below a certain point.
The fuel temperature sensing element of the air shutoff valve senses and controls fuel temperature by shifting valve the position to permit more or less air flow, thereby maintaining fuel temperature at approximately 50 deg F (10 deg C).
Should the air shutoff valve fail to close properly, resulting in fuel being heated above approximately 100 deg F (38 C), the air regulator valve will seal off ambient air pressure and route high pressure air to close the valve and seal off flow of hot air through heater.

14. Fuel Control
The engine fuel control is a hydromechanical control. Variables sensed by the fuel control are 1) power lever angle, 2) burner pressure 3) high compressor rotor bleed, 4) compressor inlet temperature.
The fuel control also utilizes these variables to control fuel flow for steady state operations, accelerations and decelerations.
The fuel control compensates automatically only for variations in altitude and airspeed when in MANUAL.
Fuel control operation may be changed from PRIMARY to MANUAL at all altitudes. If airspeed is above 225 KIAS, selection of MANUAL may be made at any throttle setting. If airspeed is below 225 KIAS, a minimum throttle setting of 65 percent rpm is required. Switch over from MANUAL to PRIMARY fuel control should be accomplished between 80 to 85 percent rpm.
The switch over to the MANUAL fuel system may be accompanied by a minor surge in engine speed and EGT.
After a switch over, the throttle should be moved slowly and smoothly to the desired power setting. Observe the engine operating limitations. When operating on the manual fuel control system, all fuel metering to the engine is accomplished by direct movement of the throttle. Power changes must be made with care, not only to prevent over speeding and extreme temperatures, but also to avoid a flameout.
Note
Complete loss of electrical power precludes switching the fuel control from the position selected.

15. ENGINE FIRE DETECTION SYSTEM
Fire Warning Light
Fire Detection System
The fire detection system will indicate the existence of fire in the area surrounding the engine, tailpipe, and accessories section. If a fire occurs in these locations, the FIRE warning light will come on. The fire detection system may be checked by depressing the master press-to-test button. When the button is depressed, the FIRE warning light will come on indicating a properly functioning circuit. The system discriminates against short circuits and prevents illumination of the fire warning light by either the fire detection control unit or press-to-test button when a short exists.
The A-4N has separated the engine/accessories section from the tail section and has two (2) fire warning lights.
The FIRE light is associated with the engine/accessories section and the FIRE TAIL is associated with the tail section.
As the temperature in the engine compartment or the fuselage aft section rises, the resistance of the sensing elements decreases. The sensing elements are connected to the control unit in a single series loop, so if an opening in the loop occurs, the numerous paths to ground permit the system to operate with no loss of efficiency. When the resistance in any part of the loop drops gradually below the set point, the fire warning circuit actuates a relay in the control unit and causes the FIRE warning light to come on.
If there is an instantaneous decrease in resistance to a point below the set point, such as that caused by a short to ground, the discriminator circuit in the control unit senses the short period of elapsed time and locks out the fire warning circuit to prevent either cockpit warning light from coming on.
TA-4
A-4N

16. Main Instrument Panel: Engine/Fuel
Engine & Fuel Instruments/ Indicators:
Located on the lower right main instrument panel (below the right HDD).
Identify the following instruments:
Exhaust Gas Temperature (EGT)
Fuel Flow
Engine RPM
Fuel Quantity
Fuel Internal/External Quantity Switch
Engine Pressure Ratio (EPR)
Oil Pressure
Oil Quantity Push-Button
TA-4
A-4N

17. Manual Fuel Light TA-4 A-4N Engine Control Panel
The engine control panel, contains all other controls for the operation of the engine. On the panel are the manual fuel control warning light, the drop tanks pressurization and flight refuel switch, the air refueling fuselage-only switch, the emergency transfer and wing fuel dump switch, the engine starter switch, and the fuel control switch.
FUEL CONTROL SWITCH. A two-position fuel control switch on the engine control panel is used to select the mode of operation of the engine fuel control unit. With the switch at PRIM, the automatic metering devices in the fuel control unit regulate the flow of fuel to the engine. The control compensates automatically for only variations in altitude and airspeed when the switch is in MANUAL.
MANUAL FUEL CONTROL WARNING LIGHT. The manual fuel control warning light, labeled FUEL CONT, on the engine control panel (TA-4) or left annunciator panel(A-4N) comes on when the fuel control has shifted to the manual mode of operation. The light indicates the position of the emergency transfer valve which directs the fuel to either the primary or manual fuel control system. The emergency transfer valve is kept in the manual fuel control position by spring load until overcome by engine-driven fuel pump pressure, regardless of the position of the fuel control switch. Consequently, the light will be on during normal engine starts until fuel pressure within the control shifts the transfer valve to the PRIMARY position at approximately 5 to 10 percent rpm. The light will also come on shortly after engine shutdown indicating a shift to the MANUAL mode upon loss of fuel pressure. Complete loss of electrical power precludes switching the fuel control from position selected.
TA-4
A-4N

18. STARTING THE ENGINE GROUND CONTROLLED ENGINE START.
1. Pilot holds one finger vertically--P/C CONNECTS EXTERNAL POWER AND STARTS GTC
2. ICS CHECK
3. Master press-to-test CHECK
4. Throttle OFF
5. Pilot holds two fingers vertically-P/C OPENS GTC AIR VALVE
6. Five-percent RPM, throttle IGN
7. Fifteen-percent RPM, throttle-----IDLE
8. At forty-five-percent RPM, three fingers held vertically------P/C CLOSE GTC AIR VALVE AND REMOVE STARTER AIR HOSE
9. Stabilized idle RPM, pilot holds four fingers vertically P/C SELECTS INTERNAL ELECTRICAL POWER, SECURES POWER UNIT, REMOVES ELECTRICAL CABLE, SECURES ACCESS PANEL
NOTE!
Lightoff should occur within 20 seconds after moving the throttle outboard to start the ignition cycle. Lightoff will be indicated by a rise in EGT after the throttle is moved to the IDLE position. Normally, the engine should be stabilized at IDLE rpm within 45 seconds after P/C opens GTC valve.
All starts will be ground controlled starts.

19. FLIGHT CREW CHECKLIST REVIEW
START MALFUNCTIONS
ENGINE FIRE (GROUND/FLIGHT)
HIGH/LOW ALTITUDE ENGINE FAILURE AND AIRSTARTS: Note the procedural differences.
CHUGS AND STALLS
OIL PRESSURE AND QUANTITY MALFUNCTIONS
FUEL CONTROL MALFUNCTION
THROTTLE LINKAGE FAILURE
CAUTION
· Before attempting restart, allow sufficient time for surplus fuel to drain from engine combustion chamber.
· Normally do not reengage starter air before complete engine rundown.

20. *1. Throttle – OFF (DO NOT ACTUATE IGNITERS)
START MALFUNCTIONS
HUNG
HOT
WET
FALSE START
*1. Throttle – OFF (DO NOT ACTUATE
IGNITERS)
*2. Engine starter – CONTINUE CRANKING.
(15-20 SECS)
3. Engine start air – CUT-OFF 1N 2N 3N 4N
ABNORMAL STARTS. The three predominant types of abnormal starts that may occur are hot start, hung start, and wet start. Do not attempt restart until the engine has stopped rotating.
WARNING
In order to preclude a hung start on P-408 engines, assure throttle is set at IDLE during start.
Hung Start. Normally, engine idle speed will be reached within one minute after throttle is advanced to the IDLE position. During a hung start, the engine lights off normally but rpm remains (hangs up) at a level below normal idle speed and EGT continues to rise toward maximum temperature. A hang-up normally occurs at approximately 40 percent RPM.
When it is evident that the engine is not going to achieve to idle rpm, or if any rise in EGT is observed retard the throttle to Off and allow the engine to rotate for 15 to 20 seconds using the starter to cool turbine section.
Probable Cause
1. Starter cutout premature.
2. Starter not accelerating engine sufficiently.
3. Incorrect calibrated fuel control.
4. Loose or broken combustion chamber static pressure sense line.
5. Engine rotor binding due to damaged bearing or damaged accessory gearbox. 5C 6N

21. To clear engine of excess fuel after abnormal starts:
CLEARING ENGINE
To clear engine of excess fuel after abnormal starts:
1. Throttle OFF (DO NOT ACTIVATE
IGNITERS)
2. Continue to crank engine with starter for 20 seconds.
3. Engine start air OFF
CAUTION
· Before attempting restart, allow sufficient time for surplus fuel to drain from engine combustion chamber.
· Normally do not reengage starter air before complete engine rundown.

22. 1N

23. ENGINE FAILURE DURING TAKEOFF
TAKEOFF EMERGENCIES
ENGINE FAILURE DURING TAKEOFF
Throttle ENSURE AT MILITARY
*1. Confirmed engine failure---ABORT
Abort impossible DECIDE TO EJECT OR
REMAIN WITH AIRCRAFT

24. LOW ALTITUDE LOSS OF THRUST/FLAMEOUT
IN-FLIGHT EMERGENCIES
LOW ALTITUDE LOSS OF THRUST/FLAMEOUT
*1. Commence ZOOM CLIMB
*2. Throttle RETARD
*3. Fuel Control MANUAL
*4. Emergency Generator EXTEND
If thrust is not regained immediately, proceed as follows:
*5. Throttle IGN then IDLE
6. Throttle ADVANCE CAUTIOUSLY and ascertain restoration of thrust
7. If engine fails to respond and
peak altitude at zoom apex is
LESS THAN 5000 FEET EJECT
If time and altitude permit:
establish glide KIAS
commence AIRSTART

25. ENGINE FAILURE
*1. Throttle RETARD
*2. Fuel Control MANUAL
If Thrust is not regained, proceed as follows:
3. Throttle OFF
4. Emergency generator EXTEND
NOTE
Ensure NORMAL/BYPASS switch NORMAL
5. Check for FIRE
IF fire exists or existed NO RESTART
IF NO FIRE; proceed as follows:
6. Descend BELOW 30,000 FT
7. Establish glide KIAS
8. Throttle IGN then IDLE
9. Monitor RPM & EGT
10. IF no relight (within 45 seconds).--IGN then IDLE

26. FUEL CONTROL MALFUNCTION
1. Throttle MATCH WITH RPM
2. Fuel control switch MANUAL
3. Throttle ADVANCE SLOWLY TO DESIRED POWER SETTING

27. ENGINE FIRE IN FLIGHT WITH OTHER INDICATIONS
*1. Throttle OFF
*2. Manual fuel control EMER OFF
shutoff control lever
*3. Emergency generator EXTEND
*4. If fire persists EJECT IMMEDIATELY
5. If fire goes out DO NOT RELIGHT
ENGINE BUT PERFORM CONTROLLED EJECTION

28. ENGINE FIRE IN FLIGHT WARNING LIGHT ONLY
*1. Throttle MIN FOR FLIGHT
2. Land as soon as possible.

29. ENGINE RPM UNTIL LANDING ASSURED OR UNTIL DECISION TO EJECT IS MADE
ABNORMAL OIL PRESSURE
1. Less than 1 minute PERMISSIBLE
2. If oil pressure remains out of limits (high or low) more than 1 Minute Maintain (80-82% A-4, % TA-4)
ENGINE RPM UNTIL LANDING ASSURED OR UNTIL DECISION TO EJECT IS MADE
3. USE PRECAUTIONARY APPROACH.

30. 1. Abort flight, land at nearest suitable landing facility.
OIL LOW LIGHT ON
1. Abort flight, land at nearest suitable landing facility.
2. Use Precautionary Approach for landing

31. THROTTLE LINKAGE FAILURE
1. Do not manipulate throttle.
2. Determine thrust available, be alert for slow RPM
decay.
3. Fly PA or STA, maintain 180 KIAS minimum.