5: Engine Systems

The Gas Turbine Engine

Types of Turbojet Engines

Engine Nomenclature

Engine Instruments

Engine Pressure Ratio (EPR)

Engine Compressor Operation

Turbine Section

Engine Fuel System

Starter

Engine Oil Systems

Normal Procedures

Abnormal Procedures

The Gas Turbine Engine

All gas turbine engines consist of an air inlet section, a compressor section, the combustion section, the turbine section and the exhaust section. The terminology used to describe gas turbines is a little inconsistent at times. Sometimes the inlet and/or the exhaust sections are considered part of the airframe rather than the engine. At a minimum, a turbine engine has a compressor, combustor and a turbine section. Engines are sometimes divided into cold and hot sections. The cold section is forward of the combustor and the hot section is from the combustor aft. See Figure 5-1.

Figure 5-1. Turbojet engine

Air enters the inlet at roughly ambient temperature and pressure. As it passes through the compressor, the pressure increases and so does the temperature due to the heat of compression. Bleed air is tapped off the compressor for such accessories as air conditioning and thermal anti-icing. Often this bleed air is taken from more than one point on the compressor or compressors.

The section connecting the compressor and the combustion sections is called the diffuser. In the diffuser, the cross-sectional area of the engine increases. This allows the air stream from the compressor to slow and its pressure to increase. In fact, the highest pressure in the engine is attained at this point.

Next, the air enters the combustion section were it is mixed with fuel and the mixture is ignited. There is no need for an ignition system that operates continuously (such as the spark plugs in a piston engine) because the uninterrupted flow of fuel and air will sustain combustion after an initial light-off. The combustion of the fuel-air mixture causes a great increase in volume, and because there is higher pressure at the diffuser, the gas exits through the turbine section. The temperature of the gas rises rapidly as it passes from the front to the rear of the combustion section. It reaches its highest point in the engine at the turbine inlet. The maximum turbine inlet temperature is a major limitation on turbojet performance and without cooling, it could easily reach up to 4,000°F, far beyond the limits of the materials used in the turbine section. To keep the temperature down to an acceptable 1,100 to 1,500°F, surplus cooling air from the compressor is mixed in aft of the burners.

The purpose of the turbine(s) is to drive the compressor(s) which are connected by a drive shaft. Since the turbines take energy from the gas, both the temperature and pressure drop.

Figure 5-2. The convergent-type jet nozzle accelerates the gases as they leave the engine.

The gases exit the turbine section at very high velocity into the tailpipe. The tailpipe is shaped so that the gas is accelerated even more reaching maximum velocity as it exits into the atmosphere. See Figure 5-2.

The four events in the energy release cycle of a gas turbine engine (intake, compression, combustion, and exhaust) are the same as in a piston engine. See Figure 5-3. The difference in the cycle is that when energy is added, the pressure remains relatively constant while the volume increases dramatically, just the reverse of a piston engine. See Figure 5-4.

Figure 5-3. A comparison between the working cycle of a gas turbine engine and a piston engine.

Figure 5-4. The Brayton cycle on a pressure-volume diagram

Types of Turbojet Engines

Turbojet engines can be divided into two types based on the type of compressor they employ. The centrifugal-flow engine uses a compressor that is very similar in design to the compressor used in piston engine turbochargers. This was the type of engine found on many early jet aircraft, but it has been almost entirely replaced by the axial-flow engine.

Just about the only contemporary use of the centrifugal-flow gas turbine engine is an APU (Auxiliary Power Unit). The main advantage of using a centrifugal-flow compressor in an APU is that it is shorter in length.

The compressor of an axial-flow engine consists of sets of rotating airfoils (rotors) alternating with sets of stationary blades (stators). Each pair of rotors and stators is a compressor stage. Modern engines can have compressors consisting of 10 or more stages.

Engine Nomenclature

FAA Figure 9 shows an exploded view of a typical turbojet engine.

Many modern engines have two independently rotating sets of compressors and turbines. There are several advantages to this type of design. These include the ability to get higher compression ratios, improved high altitude performance and greater starting flexibility.

See FAA Figure 8. Notice that the forward compressor is driven by the aft turbines and that aft compressor is driven on a concentric shaft by the forward turbine. The forward compressor is referred to as the low-pressure compressor and the turbines that drive it are the low-pressure turbines. The aft compressor is the high-pressure compressor and its turbine is the high-pressure turbine.

Engine Instruments

For the purposes of measuring internal temperatures and pressures, engine stations are established. The exact numbering system varies from engine to engine, but usually follows the same general pattern. The air inlet duct is station 1, the entrance to the first compressor stage is station 2. The tailpipe is usually station 7, 8 or 9 depending on how many compressors are used within the engine. See Figure 5-5.

Figure 5-5

The measurement taken at a station is designated by the letter P for pressure or T for temperature followed by subscripts to indicate the station and type of measurement. For example, the total pressure at station 2 is Pt2. The total temperature at station 7 is Tt7.

Compressor RPM is the measure of turbojet engine speed. A major limit on engine performance is maximum RPM and so the tachometers are monitored during flight for any possible overspeed condition. RPM response during starting is also critical. Compressor RPM is designated by the letter N. The RPM of a single-spool compressor is designated NC. On a dual-spool compressor, the RPM of the low-speed (low-pressure) compressor is N1 and the RPM of the high-speed (high-pressure) compressor is N2. See Figure 5-6.

Figure 5-6. Engine indicators

Because the actual compressor RPM will vary greatly from one type of engine to another, tachometers usually are calibrated in percent of RPM rather than the actual values. This allows various types of engines to be operated on the same basis of comparison (e.g., 100% RPM is full power in all engines).

The typical gauge uses two scales so that the RPM can be read to the nearest 1%. A large needle indicates on a scale divided into 10% increments, while a small needle uses a separate scale indicating in 1% increments. To read the RPM, add the indication of the small needle to the indication of the large needle.

RPM (particularly N1) can be used as an indicator of engine thrust output. On some engines it is the primary thrust indicator. On others it is used as a cross check against the primary indicator (EPR) during critical operations such as takeoff or go around. However, because of compressor aerodynamics, the percent of RPM is not a direct indicator of the thrust output. At lower settings, a large increase in RPM will result in a relatively small increase in thrust, and at higher settings a small change in RPM will result in a large thrust change. This non-linear response to RPM changes must be allowed for when using the tachometer to set thrust.

Engine Pressure Ratio (EPR)

Many engines use Engine Pressure Ratio (EPR) as the primary thrust indicator. Because it is much more linear than RPM and because it compensates automatically for the effects of altitude and airspeed, EPR is considered the most accurate indication of thrust in a turbojet engine.

EPR is the ratio of the total pressure between the front end of the compressor and the rear of the turbine. That is, EPR is turbine outlet pressure divided by compressor inlet pressure. For example, if the turbine discharge pressure of a turbojet engine is 62.4" Hg and the inlet pressure is 29.96" Hg, the EPR is 2.08 (62.4 ÷ 29.96 = 2.08). The EPR gauge does this calculation automatically.

The symbol for total pressure is Pt. The compressor inlet pressure is almost invariably Pt2. As mentioned earlier, the designation of turbine discharge pressure will vary depending on the type of engine. On a single-spool engine, the turbine discharge pressure will be Pt5, on a dual-spool engine it will be Pt7 and on a three-spool engine it will be Pt9. Therefore, on a dual-spool turbojet engine, EPR will equal Pt7 ÷ Pt2.

Engine Compressor Operation

As air enters the compressor section, it is directed onto the first stage rotor blades by the inlet guide vanes. These guide vanes direct the air flow onto the rotors at the proper angle of attack. Some engines don't have inlet guide vanes and must rely on duct design for proper air flow through the compressor.

The rest of the compressor consists of alternating sets of rotor and stator blades. Remember that each pair of rotors and stators constitute a stage of the compressor. Rotor and stator blades are both airfoil sections and produce lift, but they perform complementary functions within the compressor. The rotors accelerate the air rearward. The stators' function is to convert velocity energy into pressure energy, acting as divergent ducts and slowing the air back to its original velocity, which causes an increase in pressure. The stators also direct the air onto the next stage of rotors at the appropriate angle. The process is then repeated at each successive stage of the compressor. The net effect is that the air flow velocity will vary up and down around its original value as it passes through the compressor but the air pressure will constantly rise. The compressor case narrows in diameter to accommodate the reduced volume of the compressed air.

The compressor blades, particularly the rotors, can stall if the angle of attack of the air becomes excessive. One important consideration in engine design is the prevention of compressor stalls, which in extreme conditions can cause an almost immediate destruction of the engine.

Compressor angle of attack is essentially a function of compressor RPM and air flow velocity. Low air velocity tends to raise the angle of attack as do rapid engine accelerations. One of the most critical conditions occurs when the aircraft is at low speeds (or stationary) and the power is rapidly increased.

One common method of stabilizing the compressor at low and intermediate engine speeds is the use of bleed valves. These valves, which are placed at the critical point in the compressor, automatically open at low engine speed and bleed some of the compressor air overboard. This lowers the interstage pressure, which lowers the rotor blade's angle of attack.

Turbine Section

As the gases exit the combustion section they first flow through the nozzle diaphragm. This unit is also referred to as the turbine nozzle or the turbine guide vane assembly. The purpose of the nozzle diaphragm is to accelerate the hot gases to the highest possible velocity. Gas velocity can reach Mach 1 exiting the nozzle under certain conditions. A secondary function of the nozzle diaphragm is to direct the flow of gases onto the turbine buckets (blades) at the desired angle.

The function of the turbine section is to convert the kinetic and heat energy in the hot gases from the combustion section into mechanical energy that drives the compressors. Like the compressor section, the turbine section has alternating sets of stators and rotors. These blades extract energy by keeping the gases at a relatively constant velocity while allowing their pressure and temperature to drop.

The turbines are the most stressed section of the engine, having to endure exceptionally high temperatures and pressures. The turbine blades (sometimes called buckets) are usually made of super alloys, but are still very susceptible to damage if the engine operating limits are exceeded even momentarily.

Engine Fuel System

The engine fuel system usually consists of one or more engine-driven fuel pumps, a fuel heater, a fuel filter, the fuel control unit and a fuel-oil cooler.

Jet fuel often contains some entrained water. This water can damage the fuel control unit if it freezes into ice crystals and is not filtered out. A filter is usually located upstream of the fuel control unit to remove any entrained ice (and other contaminants) before the fuel enters the fuel control unit. See Figure 5-7.

Figure 5-7. Typical engine fuel system

There is a pressure-sensing element on either side of the filter and if there is a significant pressure drop across the filter, an icing or fuel filter light illuminates on the flight engineer's panel. This system does not directly sense the presence of ice in the fuel but rather that the fuel filter is becoming clogged. If nothing further is done and the filter becomes completely blocked, a filter bypass will allow fuel to flow to the engine.

When the icing light illuminates, the usual procedure is to apply fuel heat. The fuel heater most often utilizes hot bleed air from the last stage compressor or the engine diffuser. As the fuel passes through this fuel-air heat exchanger, it is warmed enough to melt any ice in the fuel and in the fuel filter which is downstream. When fuel heat is applied, there will be a slight drop in engine EPR because of the extra bleed air used for fuel heat.

Many installations have a fuel-oil cooler located downstream of the fuel filter and heater. The purpose of this heat exchanger is to cool the engine oil with fuel. When fuel heat is used, oil temperature will rise because of less efficient cooling.

Because it is possible to induce excessive oil temperatures, use of fuel heat is limited to one minute of operation. In any event, if the fuel filter bypass light remains on after using fuel heat for 1 minute, there is probably some solid contamination other than ice in the fuel filter.

If the fuel temperature is below 32°F, fuel heat should be used as a precaution for 1 minute prior to takeoff or landing and for 1 minute every 30 minutes in flight. Because of the degradation of engine performance due to heat damage to the fuel control or vapor lock, fuel heat should not be used during takeoff, approach to landing, or go-arounds.

Starter

The starter on a gas turbine engine is used to rotate the compressor to an RPM where the air flow through the engine is sufficient to sustain combustion. When this RPM is reached, an ignition source is activated and fuel is added. When light-off occurs, the starter usually remains engaged to aid the engine's acceleration toward idle power. See Figure 5-8.

On dual-compressor engines, only the high-speed compressor (N2) is rotated by the starter. Just prior to reaching minimum starting RPM, the air flow through the engine is enough to cause the rotation of the N1 compressor.

Because of its superior power-to-weight ratio, the air turbine starter also called pneumatic starter, is the one most commonly found on larger turbojet engines. This starter uses high-pressure air to drive a small turbine, which in turn mechanically drives the compressor section. The high-pressure air usually reaches the starter through the aircraft pneumatic system and can be supplied by an auxiliary power unit, another running engine or a ground power unit. The air is vented overboard after driving the starter turbine.

The air turbine part of the starter turns at very high RPM and does not have to go very much above its normal operating speed to burst with catastrophic results for the starter and engine. To protect against this possibility, most starters are equipped with a clutch assembly to disengage the starter from the compressor drive shaft if the RPM becomes excessive.

Pneumatic starters have very limited cooling and can overheat if run too long. A duty cycle limit will be prescribed in the airplane flight manual.

Figure 5-8. Air turbine starter

Engine Oil Systems

The oil system is designed to cool and lubricate those parts of the engine which are subject to high friction loads from engine rotation and to high temperatures from combustion. The turbine bearings are the units from which the oil system extracts the most heat. In fact, cooling these bearings is so critical that some engines augment the oil cooling with bleed air directed onto the turbine wheel.

The two types of oil systems found in turbine engines are the dry sump and the wet sump. The main difference is that the wet sump system has an oil tank which is integral within the engine while the dry sump usually has an externally mounted tank. Almost all oil systems on modern turbine engines are of the dry sump variety.

The typical dry sump system consists of an oil tank, an engine-driven oil pump, the main oil filter, an oil cooler, temperature and pressure sensors, a last chance filter and scavenge lines and pumps to return the oil to the tank. These components are divided into the pressure, breather and scavenge subsystems. Since these systems usually spray oil onto the bearings, the terms dry sump, pressure and spray describe them. See Figure 5-9.

A filter is installed downstream of the tank to remove any particles that may be suspended in the oil. This filter is equipped with a bypass so that if it becomes clogged, oil can still flow to the engine. This condition cannot be tolerated for long since the unfiltered oil can cause damage to the engine bearings and possibly clog the last chance filters which would deny oil to the engine. Typically, there is a pressure sensor on either side of the filter. If there is an impending filter bypass, the pressure drop across the filter will cause a warning light to illuminate.

The engine oil is cooled in a liquid-to-liquid heat exchanger by fuel flowing to the engine. A thermostatic controlling device causes the oil to bypass the cooling elements when its temperature is below the normal operating range. As the oil is heated, more oil is routed through the cooler to maintain the oil temperature within limits.

Viscosity of oil is a measurement of the oil's thickness. It is determined by measuring the time required for 60 cubic centimeters of oil to pass through a calibrated orifice. The viscosity of oil will vary with temperature, growing less at higher temperatures. The Oil Viscosity Index is based on the rate of change in oil viscosity relative to temperature.

Figure 5-9. Typical engine dry sump oil system

Normal Procedures

To start a turbojet engine, you need low-pressure air (around 30 PSI) and a source of electrical power. The air rotates the engine's starter and the electrical power is used to fire the igniters. The air pressure can be supplied by an auxiliary power unit (APU), another running engine or a ground power unit. The air is supplied to the engine starter through the aircraft's normal pneumatic system. The air pressure before, during and after start can be monitored on the pneumatic system air pressure gauge.

Electrical power can be supplied from an APU, an electrical ground power unit, an aircraft generator driven by a running engine or, in some aircraft, the battery. Besides powering the igniters, the electrical system will power some of the engine instruments (i.e., EPR, fuel flow, oil pressure and pneumatic system air pressure). If making a start using the battery for electrical power, only the RPM and EGT gauges will operate.

The proper start sequence for any turbojet engine is starter, ignition, then fuel. That is, the starter is engaged and the compressor is rotated to some minimum RPM. When the start RPM is attained, the ignition system is activated and then fuel is introduced into the burner section. Engine light-off should occur almost immediately—certainly within 20 seconds. This is indicated by a rise in EGT followed by accelerating engine RPM. The starter remains engaged to help the engine accelerate to idle RPM. At the appropriate RPM (idle or just below) the starter is disengaged and the ignition is shut off. Turbojet engines will idle at from 40 to 60% RPM (N2).

The most critical parameter during engine start is EGT or TIT. If the starting temperature limits are exceeded even for a few seconds, the engine must be inspected or even removed and repaired. For this reason maximum temperature must be noted on each start. A typical start procedure is as follows:

1. Ensure that electrical power is available to the ignition system.

2. Ensure that the minimum pneumatic duct pressure is available to the starter.

3. Move the start switch to "ground start" (this opens the starter air valve and arms the ignition).

4. Confirm the start valve opening by a drop in pneumatic duct pressure followed by increasing RPM.

5. At the minimum starting RPM, move the start lever to "idle" (this activates the ignition and opens the fuel valve in the fuel control unit).

6. Monitor EGT for a rapid rise in temperature (confirms engine light-off).

7. Monitor EGT, RPM, oil pressure and fuel flow for possible malfunctions.

8. At the appropriate RPM, move the start switch to "off" (closes the starter air valve and deactivates the ignition).

9. Confirm starter air valve closure by a rise in pneumatic system duct pressure.

10. Monitor EGT, RPM, oil pressure and fuel flow for proper values at idle power.

An engine is shutdown by moving the start lever to cutoff. This cuts off fuel at the fuel control unit. Turbine wheel cooling is very critical, so the engine should be run at idle power for at least one minute before shutdown. If this is not done, the turbine case could shrink down enough to seize the turbine blades.

Abnormal Procedures

A hot start occurs when engine EGT exceeds the limits set by the manufacturer during the start cycle. A hot start can be anticipated if the fuel flow during start exceeds the normal range. A hung start or false start occurs when the engine lights-off normally but does not accelerate to idle RPM. EGT initially remains within normal limits but it will tend to rise slowly above red line if nothing is done. A hung start often results when the starter cuts out below the RPM at which the engine will accelerate to idle on its own.

A hot start (including high fuel flow before light-off), a hung start and an engine fire on start are all dealt with the same way. Fuel and ignition should be shut off and the starter should remain engaged for a specified time (usually one minute). Shutting off fuel and ignition will prevent or at least minimize engine damage from excessive temperatures. Motoring the engine with the starter will cool the engine and blow any residual fuel out the tailpipe. The cause of any abnormal start should be investigated before attempting a second start, even if no engine limits were exceeded.

The oil pump is engine driven and there should be an indication of oil pressure before engine light-off. If there isn't, discontinue the start and investigate the cause.

High EGT at takeoff power settings is usually caused by a compressor bleed valve malfunction that reduces cooling air flow through the burner section.

If the Pt2 probe in the engine nose dome becomes iced over, false EPR indications will result. The most unusual result is an erroneously high reading. This occurs because the Pt2 pressure is lower than normal while the exhaust pressure is unaffected. The greater-than-normal difference in the two values causes a high reading. Setting a computed takeoff EPR when this condition exists will result in the engine producing less than takeoff power.

A falsely low EPR reading can occur if the engine anti-ice is turned on after the Pt2 probe has iced over. The anti-ice air will pressurize the nose dome driving the Pt2 pressure too high and causing the EPR to be too low.

Takeoff power should be set using the computed EPR for the runway conditions. With the EPR set, engine RPM should be cross-checked to ensure takeoff power is actually set.

High engine temperatures (EGT or TIT) in cruise flight can indicate damage to the compressor section. Significant amounts of airborne particles can form enough of a coating on the compressor blades to cause higher than normal EGT readings.

The turbine blades are particularly susceptible to damage from excessive temperature. Creep is the elongation of turbine rotor blades due to high torsion and heat stresses. A certain amount of turbine blade creep is inevitable but temperatures above the specified limits can use up the turbine life very rapidly. Stress ruptures of turbine blades can also occur in high temperature situations.

Indications of turbine damage or loss of turbine efficiency are high EGT, high fuel flow and low engine RPM at all power settings.