Turbine engines produce thrust by increasing the velocity of the air flowing through the engine. A turbine engine consists of an air inlet, compressor, combustion chambers, turbine section, and exhaust.
Figure 1: Basic components of a turbine engine.
The turbine engine has the following advantages over a reciprocating engine: less vibration, increased aircraft performance, reliability, and ease of operation.
Types of turbine engines
Turbine engines are classified according to the type of compressors they use. The compressor types fall into three categories—centrifugal flow, axial flow, and centrifugal-axial flow. Compression of inlet air is achieved in a centrifugal flow engine by accelerating air outward perpendicular to the longitudinal axis of the machine. The axial-flow engine compresses air by a series of rotating and stationary airfoils moving the air parallel to the longitudinal axis. The centrifugalaxial flow design uses both kinds of compressors to achieve the desired compression.
The path the air takes through the engine and how power is produced determines the type of engine. There are four types of aircraft turbine engines—turbojet, turboprop, turbofan, and turboshaft.
The turbojet engine contains four sections: compressor, combustion chamber, turbine section, and exhaust. The compressor section passes inlet air at a high rate of speed to the combustion chamber. The combustion chamber contains the fuel inlet and igniter for combustion. The expanding air drives a turbine, which is connected by a shaft to the compressor, sustaining engine operation. The accelerated exhaust gases from the engine provide thrust. This is a basic application of compressing air, igniting the fuel-air mixture, producing power to self-sustain the engine operation, and exhaust for propulsion.
Turbojet engines are limited on range and endurance. They are also slow to respond to throttle applications at slow compressor speeds.
A turboprop engine is a turbine engine that drives a propeller through a reduction gear. The exhaust gases drive a power turbine connected by a shaft that drives the reduction gear assembly. Reduction gearing is necessary in turboprop engines because optimum propeller performance is achieved at much slower speeds than the engine’s operating r.p.m. Turboprop engines are a compromise between turbojet engines and reciprocating powerplants. Turboprop engines are most efficient at speeds between 250 and 400 m.p.h. and altitudes between 18,000 and 30,000 feet. They also perform well at the slow airspeeds required for takeoff and landing, and are fuel efficient. The minimum specific fuel consumption of the turboprop engine is normally available in the altitude range of 25,000 feet to the tropopause.
Turbofans were developed to combine some of the best features of the turbojet and the turboprop. Turbofan engines are designed to create additional thrust by diverting a secondary airflow around the combustion chamber. The turbofan bypass air generates increased thrust, cools the engine, and aids in exhaust noise suppression. This provides turbojet-type cruise speed and lower fuel consumption.
The inlet air that passes through a turbofan engine is usually divided into two separate streams of air. One stream passes through the engine core, while a second stream bypasses the engine core. It is this bypass stream of air that is responsible for the term “bypass engine.” A turbofan’s bypass ratio refers to the ratio of the mass airflow that passes through the fan divided by the mass airflow that passes through the engine core.
The fourth common type of jet engine is the turboshaft.
It delivers power to a shaft that drives something other than a propeller. The biggest difference between a turbojet and turboshaft engine is that on a turboshaft engine, most of the energy produced by the expanding gases is used to drive a turbine rather than produce thrust. Many helicopters use a turboshaft gas turbine engine. In addition, turboshaft engines are widely used as auxiliary power units on large aircraft.
It is possible to compare the performance of a reciprocating powerplant and different types of turbine engines. However, for the comparison to be accurate, thrust horsepower (usable horsepower) for the reciprocating powerplant must be used rather than brake horsepower, and net thrust must be used for the turbine-powered engines. In addition, aircraft design configuration, and size must be approximately the same.
BHP Brake horsepower is the horsepower actually delivered to the output shaft. Brake horsepower is the actual usable horsepower.
Net Thrust The thrust produced by a turbojet or turbofan engine.
THP Thrust horsepower is the horsepower equivalent of the thrust produced by a turbojet or turbofan engine.
ESHP Equivalent shaft horsepower, with respect to turboprop engines, is the sum of the shaft horsepower (SHP) delivered to the propeller and the thrust horsepower (THP) produced by the exhaust gases.
Figure 2: Engine net thrust versus aircraft speed and drag.
Figure 2 shows how four types of engines compare in net thrust as airspeed is increased. This figure is for explanatory purposes only and is not for specific models of engines. The four types of engines are:
- Reciprocating powerplant.
- Turbine, propeller combination (turboprop).
- Turbine engine incorporating a fan (turbofan).
- Turbojet (pure jet).
The comparison is made by plotting the performance curve for each engine, which shows how maximum aircraft speed varies with the type of engine used. Since the graph is only a means of comparison, numerical values for net thrust, aircraft speed, and drag are not included.
Comparison of the four powerplants on the basis of net thrust makes certain performance capabilities evident.
In the speed range shown to the left of Line A, the reciprocating powerplant outperforms the other three types. The turboprop outperforms the turbofan in the range to the left of Line C. The turbofan engine outperforms the turbojet in the range to the left of Line F. The turbofan engine outperforms the reciprocating powerplant to the right of Line B and the turboprop to the right of Line C. The turbojet outperforms the reciprocating powerplant to the right of Line D, the turboprop to the right of Line E, and the turbofan to the right of Line F.
The points where the aircraft drag curve intersects the net thrust curves are the maximum aircraft speeds. The vertical lines from each of the points to the baseline of the graph indicate that the turbojet aircraft can attain a higher maximum speed than aircraft equipped with the other types of engines. Aircraft equipped with the turbofan engine will attain a higher maximum speed than aircraft equipped with a turboprop or reciprocating powerplant.
Turbine engine instruments
Engine instruments that indicate oil pressure, oil temperature, engine speed, exhaust gas temperature, and fuel flow are common to both turbine and reciprocating engines. However, there are some instruments that are unique to turbine engines. These instruments provide indications of engine pressure ratio, turbine discharge pressure, and torque. In addition, most gas turbine engines have multiple temperature-sensing instruments, called thermocouples, that provide pilots with temperature readings in and around the turbine section.
Engine pressure ratio
An engine pressure ratio (EPR) gauge is used to indicate the power output of a turbojet/turbofan engine.
EPR is the ratio of turbine discharge to compressor inlet pressure. Pressure measurements are recorded by probes installed in the engine inlet and at the exhaust.
Once collected, the data is sent to a differential pressure transducer, which is indicated on a cockpit EPR gauge.
EPR system design automatically compensates for the effects of airspeed and altitude. However, changes in ambient temperature do require a correction to be applied to EPR indications to provide accurate engine power settings.
Exhaust gas temperature
A limiting factor in a gas turbine engine is the temperature of the turbine section. The temperature of a turbine section must be monitored closely to prevent overheating the turbine blades and other exhaust section components. One common way of monitoring the temperature of a turbine section is with an exhaust gas temperature (EGT) gauge. EGT is an engine operating limit used to monitor overall engine operating conditions.
Variations of EGT systems bear different names based on the location of the temperature sensors. Common turbine temperature sensing gauges include the turbine inlet temperature (TIT) gauge, turbine outlet temperature (TOT) gauge, interstage turbine temperature (ITT) gauge, and turbine gas temperature (TGT) gauge.
Turboprop/turboshaft engine power output is measured by the torquemeter. Torque is a twisting force applied to a shaft. The torquemeter measures power applied to the shaft. Turboprop and turboshaft engines are designed to produce torque for driving a propeller.
Torquemeters are calibrated in percentage units, foot-pounds, or pounds per square inch.
N1 represents the rotational speed of the low pressure compressor and is presented on the indicator as a percentage of design r.p.m. After start the speed of the low pressure compressor is governed by the N1 turbine wheel. The N1 turbine wheel is connected to the low pressure compressor through a concentric shaft.
N2 represents the rotational speed of the high pressure compressor and is presented on the indicator as a percentage of design r.p.m. The high pressure compressor is governed by the N2 turbine wheel. The N2 turbine wheel is connected to the high pressure compressor through a concentric shaft.
Figure 3: Dual-spool axial-flow compressor.
Turbine engine operational considerations
Because of the great variety of turbine engines, it is impractical to cover specific operational procedures.
However, there are certain operational considerations that are common to all turbine engines. They are engine temperature limits, foreign object damage, hot start, compressor stall, and flameout.
Engine temperature limitations
The highest temperature in any turbine engine occurs at the turbine inlet. Turbine inlet temperature is therefore usually the limiting factor in turbine engine operation.
Turbine engine thrust varies directly with air density.
As air density decreases, so does thrust. While both turbine and reciprocating powered engines are affected to some degree by high relative humidity, turbine engines will experience a negligible loss of thrust, while reciprocating engines a significant loss of brake horsepower.
Foreign object damage
Due to the design and function of a turbine engine’s air inlet, the possibility of ingestion of debris always exists. This causes significant damage, particularly to the compressor and turbine sections. When this occurs, it is called foreign object damage (FOD). Typical FOD consists of small nicks and dents caused by ingestion of small objects from the ramp, taxiway, or runway.
However, FOD damage caused by bird strikes or ice ingestion can also occur, and may result in total destruction of an engine.
Prevention of FOD is a high priority. Some engine inlets have a tendency to form a vortex between the ground and the inlet during ground operations. A vortex dissipater may be installed on these engines.
Other devices, such as screens and/or deflectors, may also be utilized. Preflight procedures include a visual inspection for any sign of FOD.
Turbine engine hot/hung start
A hot start is when the EGT exceeds the safe limit. Hot starts are caused by too much fuel entering the combustion chamber, or insufficient turbine r.p.m. Any time an engine has a hot start, refer to the AFM, POH, or an appropriate maintenance manual for inspection requirements.
If the engine fails to accelerate to the proper speed after ignition or does not accelerate to idle r.p.m., a hung start has occurred. A hung start, may also be called a false start. A hung start may be caused by an insufficient starting power source or fuel control malfunction.
Compressor blades are small airfoils and are subject to the same aerodynamic principles that apply to any airfoil. A compressor blade has an angle of attack. The angle of attack is a result of inlet air velocity and the compressor’s rotational velocity. These two forces combine to form a vector, which defines the airfoil’s actual angle of attack to the approaching inlet air.
A compressor stall can be described as an imbalance between the two vector quantities, inlet velocity and compressor rotational speed. Compressor stalls occur when the compressor blades’ angle of attack exceeds the critical angle of attack. At this point, smooth airflow is interrupted and turbulence is created with pressure fluctuations. Compressor stalls cause air flowing in the compressor to slow down and stagnate, sometimes reversing direction.
Figure 4: Comparison of normal and distorted airflow into the compressor section.
Compressor stalls can be transient and intermittent or steady state and severe. Indications of a transient/intermittent stall are usually an intermittent “bang” as backfire and flow reversal take place. If the stall develops and becomes steady, strong vibration and a loud roar may develop from the continuous flow reversal. Quite often the cockpit gauges will not show a mild or transient stall, but will indicate a developed stall. Typical instrument indications include fluctuations in r.p.m., and an increase in exhaust gas temperature. Most transient stalls are not harmful to the engine and often correct themselves after one or two pulsations. The possibility of engine damage, which may be severe, from a steady state stall is immediate.
Recovery must be accomplished quickly by reducing power, decreasing the airplane’s angle of attack and increasing airspeed.
Although all gas turbine engines are subject to compressor stalls, most models have systems that inhibit these stalls. One such system uses variable inlet guide vane (VIGV) and variable stator vanes, which direct the incoming air into the rotor blades at an appropriate angle. The main way to prevent air pressure stalls is to operate the airplane within the parameters established by the manufacturer. If a compressor stall does develop, follow the procedures recommended in the AFM or POH.
A flameout is a condition in the operation of a gas turbine engine in which the fire in the engine unintentionally goes out. If the rich limit of the fuel/air ratio is exceeded in the combustion chamber, the flame will blow out. This condition is often referred to as a rich flameout. It generally results from very fast engine acceleration, where an overly rich mixture causes the fuel temperature to drop below the combustion temperature. It also may be caused by insufficient airflow to support combustion.
Another, more common flameout occurrence is due to low fuel pressure and low engine speeds, which typically are associated with high-altitude flight. This situation also may occur with the engine throttled back during a descent, which can set up the lean-condition flameout. A weak mixture can easily cause the flame to die out, even with a normal airflow through the engine.
Any interruption of the fuel supply also can result in a flameout. This may be due to prolonged unusual attitudes, a malfunctioning fuel control system, turbulence, icing or running out of fuel.
Symptoms of a flameout normally are the same as those following an engine failure. If the flameout is due to a transitory condition, such as an imbalance between fuel flow and engine speed, an airstart may be attempted once the condition is corrected. In any case, pilots must follow the applicable emergency procedures outlined in the AFM or POH. Generally, these procedures contain recommendations concerning altitude and airspeed where the airstart is most likely to be successful.
This concludes the excursion to turbine engines. Please return to the Aircraft Powerplant page to continue study.