Unit Specific 202
Propulsion Systems

The following system components and component parts are discussed below mentioning their function, location, and safety/protective devices.

202.1.1 Engine:

1. Power section- consists of an axial-flow compressor, a combustion chamber, a multi-stage turbine, and an exhaust section. The last two stages of the turbine are used to drive the propeller using the torquemeter assembly and the reduction gear assembly.
   

2. Torque meter- electronically measures the torsional deflection (twist). Torsional deflection occurs in the power transmitting shaft that connects the power section to the reduction gear assembly. This torsional deflection is recorded as horsepower.
   
3. Reduction Gear system- reduces the engine rpm within the range of efficient propeller rpm. The ratio on some installations is as high as 12 or 13 to 1. This large reduction ratio is necessary because the gas turbine must operate at a very high rpm to produce power efficiently. This engine operates at a constant rpm. The propeller blade angle changes for an increase or decrease in power while the engine rpm’s remain the same.

 202.1.2 Propeller blades

The four-bladed Hamilton Standard (54H60-77) propeller provides an efficient and flexible means of converging engine SHP to thrust. The propeller consists of two principal sections: the rotating section comprises the blades, hub, spinner, and the dome that houses the pitch changing mechanism; the nonrotating section contains an oil reservoir, pressure and scavenge pumps, the governor, and control mechanism. It is a constant-speed, variable-pitch, full-feathering propeller, having the added features of negative torque sensing, pitchlock (to prevent excessive overspeed), and a combination synchronizing and synchrophasing system. The explaination of the propeller nomenclature is listed below:

5(mod #)4(#blades)H(shank size/type)60(spline size)77(minor mod#)

202.1.3 Auxiliary Power Unit

The APU is what makes the P-3 a self-sustaining aircraft. The APU is made up of a turbine compressor driving a generator that is identical to the engine-driven generators. The gas turbine compressor has a two-stage centrifugal compressor and a single-stage inward flow radial turbine. Air bled from the compressor is used for engine starting, ground air-conditioning, or for bomb bay heating. Because the power developed by the APU is somewhat limited, all of the features cannot be used simultaneously. If bleed air is demanded in sufficient quantities to jeopardize the generator output, the amount of bleed air being delivered is automatically reduced. Ground air-conditioning and engine-starting air cannot be used simultaneously. It is possible, and permissible, to use ground air and bomb bay heating simultaneously. The APU can be operated in flight for electrical power use, but bleed air is not available.

On some aircraft the GTCP 95-3 APU is installed that produces an increased airflow from the GTC 95-2 model. Though both models are interchangeable, they do not have the same EGT limitations.

202.1.4 Fuel system:

1. Fuel cells- four integral wing tanks and an auxiliary tank carry the fuel supply for the engines. The auxiliary tank, identified as tank No. 5, consists of a bladder-type fuselage tank connected to an integral center-section tank. The bladder cell is located in the unpressurized area of the lower fuselage forward of the integral tank. All tanks are automatically protected against excessive positive and negative pressure during fueling, transfer, and defueling.
   

2. Vents- the tanks are vented by float-type vent valves, located one in each wing tank and one in each cell of tank No. 5, that prevent overpressurization and overflow or siphoning during maneuvering.
   
3. Fuel boost pumps- each wing tank is equipped with a fuel boost pump consisting of a scavenge section and boost section. The scavenge section routes fuel into a surge box while the boost section pumps fuel from the surge box to the engine-driven pump. In addition, they supply fuel flow for crossfeeding. Normal boost pump pressure is 15-30 psi. A thermal switch disconnects a transfer pump whenever the case temperature of the pump exceeds 400° F.
   
4. Fuel transfer pumps- the bladder cell is equipped with two transfer pumps. Each transfer pump consists of a scavenge section and a boost section. The boost section pumps fuel from the fuselage cell to any or all wing tanks. The scavenge section pumps fuel to the fuselage cell from the bottom of the center section tank. A thermal switch disconnects a transfer pump whenever the case temperature of the pump exceeds 400° F.
   
5. Explosion suppressant foam- aircraft incorporating AFC-517 have explosion suppressant foam installed in the four integral wing tanks. The fuel cell foam is a fully reticulated fire screen designed to prevent fuel tank explosions caused by tracers or high explosive incendiary rounds, thereby igniting oxygen-rich fuel vapors. The foam adheres to fuel droplets in order to deep the fuel cell cavity too fuel rich to support combustion.
   
6. Control panels- the fueling control panel, located between the pressure fuel adapters, is the control center for fueling and defueling. Also, the panel enables the pressure fueling valves to be checked before fueling begins. Service lights on the fueling control panel will illuminate whenever the fueling panel service door is opened. An inclinometer is mounted adjacent to the panel for attitude reference when using the hydrostatic fuel quantity gauge equipment.

202.2.1 How do the following components work together to achieve the system’s function:

1. Power section- consists of a 14-stage, axial-flow compressor; six cylindrical combustion liners that comprise the combustion section; a 4-stage turbine section; an accessory drive unit; an oil system; and a fuel control unit.
   

2. Torquemeter- the struts and the torquemeter housing rigidly connect the reduction gear and the power unit. The torquemeter transmits torque from the power unit to the reduction gear assembly and provides an accurate means of measuring this torque. The torquemeter housing is also the primary support structure between the power unit and the reduction gearbox.
   
3. Reduction Gear system- reduces the high-RPM, low-torque output of the power section to a low-RPM, high-torque output to be utilized by the propeller shaft. The reduction gearbox consists of two stages of reduction to avoid excessive gearbox size. The first stage (spur gear) has a reduction ratio of 3.125:1; the second stage (planetary type) has a reduction ratio of 4.333:1. This results in an overall reduction of 13.54:1.
   
4. Propeller- the purpose of the constant speed propeller is to maintain a pre-selected RPM automatically.
   
5. Fuel cells/tanks- four integral wing tanks and an auxiliary tank carry the fuel supply for the engines.
   
6. Vents- the tanks are vented by float-type vent valves, located one in each wing tank and one in each cell of tank No. 5, that prevent overpressurization and overflow or siphoning during maneuvering.
   
7. Pumps- each wing tank is equipped with a fuel boost pump consisting of a scavenge section and boost section. The scavenge section discharges into a surge box while the boost section pumps fuel from the surge box to the engine-driven pump. In addition, they supply fuel flow for crossfeeding. Normal boost pump pressure is 15 to 30 psi.
   
8. Control panel- the control center for fueling and defueling.

202.3.1 What is the P-3C total fuel capacity in U.S. gallons?

9,200 gallons total

JP-4 = 59,800 lbs. @ 6.5

JP-5 = 62,560 lbs. @ 6.8

JP-8 = 61,640 lbs. @ 6.7

202.4.1 How does the ambient air temperature influence the operation of the Fuel system?

Fuel quantity indication can vary even though the aircraft is serviced with the same number of gallons of fuel. The factors that cause the fuel weight to change with a constant quantity are temperature and fuel density tolerances. Fuel production specifications for JP-4 and JP-5 permit a density range of ± 0.2 pounds per U.S. gallon. JP-8 fuel specifications allow a density range of ± 0.25 pounds per U.S. gallon.

For example, although JP-4 has a nominal fuel density of 6.5 pounds per U.S. gallon at 15° C, the same fuel at a temperature of 40° C has a density of 6.15 pounds per U.S. gallon. For an aircraft with 9,200 gallons of fuel, the load would be 56,580 pounds as compared with 58,510 pounds for nominal JP-4 at the same temperature.

202.4.2 How does the Fuel System interface with the following:

1. Propulsion system- each of the four wing tanks can supply fuel to its respective engine or fuel can be supplied from any tank to any engine through a crossfeed system.
   

2. Hydraulic system- the No. 2 and No. 3 fuel tanks provides a means of cooling for hydraulic pumps. Minimum of 1,000 lbs of fuel required in each tank to provide adequate hydraulic cooling.

202.5.1 What safety precautions must be observed during fueling operations?

• Any RF transmission is a potential source of fuel ignition. Use of transmitting equipment during fueling operations should be avoided.
  

• Allowing at least 3 minutes following refueling before using the dipstick. Failure to do so may result in static discharge.
  
• Aircraft incorporating AFC-517 build up and retain electrostatic charges during fueling/defueling operations. Following fueling/defueling of aircraft with AFC 517 installed, allow at least 15 minutes to dissipate electrostatic charges prior to using the dipstick or hydrostatic fuel gauge. The dipstick or hydrostatic fuel gauge shall not be used when the aircraft is exposed to temperatures below –34 degrees C. Overwing gravity fueling operations shall not be performed on aircraft incorporating AFC-517 when outside air temperature is below –6 degrees C.
  
• Fueling must be halted immediately if during the pressure fueling cycle any of the following occurs: 1) any wing tank or tank 5 is overfilled, 2) wing tank fuel spills from a wingtip vent, or 3) loud or unusual noise is accompanied by wing vibration or aircraft decking vibration. An inspection of the internal wing structure (WS 380 through WSA465) for structural damage must be accomplished prior to the next flight.

• To prevent structural damage, verify positive fuel tank venting and ensure that fueling pressure does not exceed 55 psi. The pressure gauge for tank No. 5 must be closely monitored during fueling to prevent tank over-pressurization.

• Because of wing dihedral, the filler wells must be located near the high outboard end of the engine feed tanks. Consequently, dipsticking through the filler wells will yield a no-reading indication unless the inboard tanks contain 4,225 pounds (650 gallons) of fuel or more, and the outboard tanks contain 8,775 pounds (1,350 gallons) of fuel or more.