Turboprop, Turboshaft, and Engine Performance

1-55 Spacers Seal housing SealsBearing inner race basic gas turbine to the turboprop. The typical turboprop engine can be broken down into assemblies as follows: 1. The power section assembly—contains the usual major components of a gas turbine engine (i.e., compressor, combustion chamber, turbine, and exhaust sections). 2. The reduction gear or gearbox assembly—contains those sections unique to turboprop configurations. 3. The torquemeter assembly—transmits the torque from the engine to the gearbox of the reduction section. 4. The accessory drive housing assembly—mounted on the bottom of the compressor air inlet housing.

It includes the necessary gear trains for driving all power section driven accessories at their proper rpm in relation to engine rpm.

Turboshaft Engines

A gas-turbine engine that delivers power through a shaft to operate something other than a propeller is referred to as a turboshaft engine. [Figure 1-79] The output shaft may be coupled directly to the engine turbine, or the shaft may be driven by a turbine of its own (free turbine) located in the exhaust stream. As mentioned with the turboprop, the free turbine rotates independently. This principle is used extensively in current production of turboshaft engines. The turboshaft engine’s output is measured in horsepower instead of thrust because the power output is a turning shaft. Turbofan Engines The turbofan gas turbine engine is, in principle, the same as a turboprop, except that the propeller is replaced by a duct- enclosed axial-flow fan. [Figure 1-80] The fan can be a part of the first-stage compressor blades or can be mounted as a separate set of fan blades. The blades are mounted forward of the compressor.

The general principle of the fan engine is to convert more of the fuel energy into pressure. With more of the energy 1-56 converted to pressure, a greater product of pressure times area can be achieved. One of the major advantages is turbofan production of this additional thrust without increasing fuel flow. The end result is fuel economy with the consequent increase in range. Because more of the fuel energy is turned into pressure in the turbofan engine, additional stages must be added in the turbine section to provide the power to drive the fan. This means there is less energy left over and less thrust from the core exhaust gases. Also, in a mixed-exhaust nozzle (where fan air and core air mix in a common nozzle before entering ambient conditions) the exhaust nozzle must be larger in area. The result is that the fan develops most of the thrust. The thrust produced by the fan more than makes up for the decrease in thrust of the core (gas generator) of the engine. Depending on the fan design and bypass ratio, it produces 80 percent of the turbofan engine’s total thrust.

Two different exhaust nozzle designs are used with turbofan engines. The air leaving the fan can be ducted overboard by a separate fan nozzle [Figure 1-43], or it can be ducted along the outer case of the basic engine to be discharged through the mixed nozzle (core and fan exhaust together). The fan air is either mixed with the exhaust gases before it is discharged (mixed or common nozzle), or it passes directly to the atmosphere without prior mixing (separate nozzle). Turbofans are the most widely used gas turbine engine for air transport aircraft. The turbofan is a compromise between the good operating efficiency and high thrust capability of a turboprop and the high speed, high altitude capability of a turbojet.

Turbine Engine Operating Principles The principle used by a gas turbine engine as it provides force to move an airplane is based on Newton’s Third Law. This law states that for every action there is an equal and opposite reaction; therefore, if the engine accelerates a mass of air (action), it applies a force on the aircraft (reaction). The turbofan generates thrust by giving a relatively slower acceleration to a large quantity of air. The old pure turbojet engine achieves thrust by imparting greater acceleration to a smaller quantity of air. This was its main problem with fuel consumption and noise.

The mass of air is accelerated within the engine by the use of a continuous-flow cycle. Ambient air enters the inlet diffuser where it is subjected to changes in temperature, pressure, and velocity due to ram effect. The compressor then increases pressure and temperature of the air mechanically. The air continues at constant pressure to the burner section where its temperature is increased by combustion of fuel. The energy is taken from the hot gas by expanding through a turbine which drives the compressor, and by expanding through an exhaust nozzle designed to discharge the exhaust gas at high velocity to produce thrust.

1-57 The high velocity gases from the engine may be considered continuous, imparting this force against the aircraft in which it is installed, thereby producing thrust. The formula for thrust can be derived from Newton’s second law, which states that force is proportional to the product of mass and acceleration. This law is expressed in the following formula: In the above formula, mass is similar to weight, but it is actually a different quantity. Mass refers to the quantity of matter, while weight refers to the pull of gravity on that quantity of matter. At sea level under standard conditions, 1 pound of mass has a weight of 1 pound. To calculate the acceleration of a given mass, the gravitational constant is used as a unit of comparison. The force of gravity is 32.2 feet per second squared (ft/sec 2). This means that a free falling 1 pound object accelerates at the rate of 32.2 feet per second each second that gravity acts on it. Since the object mass weighs 1 pound, which is also the actual force imparted to it by gravity, it can be assumed that a force of 1 pound accelerates a 1 pound object at the rate of 32.2 ft/sec2.

F = M x A where: M = mass in pounds per second F = force in pounds A = acceleration in feet per second Also, a force of 10 pound accelerates a mass of 10 pound at the rate of 32.2 ft/sec2. This is assuming there is no friction or other resistance to overcome. It is now apparent that the ratio of the force (in pounds) is to the mass (in pounds) as the acceleration in ft/sec 2 is to 32.2. Using M to represent the mass in pounds, the formula may be expressed thus: In any formula involving work, the time factor must be considered. It is convenient to have all time factors in equivalent units (i.e., seconds, minutes, or hours). In calculating jet thrust, the term “pounds of air per second” is convenient, since the second is the same unit of time used for the force of gravity.

F = or where: M = mass F = force A = acceleration F M A G = MA G G = gravity 1-58 Thrust Using the following formula, compute the force necessary to accelerate a mass of 50 pounds by 100 ft/sec2. F = MA G F = 50 lb x 100 ft/sec2 32.2 ft/sec2 F = 5,000 lb-ft/sec2 32.2 ft/sec2 F = 155 lb This illustrates that if the velocity mass per second is increased by 100, the resulting thrust is 155 pounds. Since the turbojet engine accelerates air, the following formula can be used to determine jet thrust: As an example, to use the formula for changing the velocity of 100 pounds of mass airflow per second from 600 ft/sec to 800 ft/sec, the formula can be applied as follows: As shown by the formula, if the mass airflow per second and the difference in the velocity of the air from the intake to the exhaust are known, it is easy to compute the force necessary to produce the change in the velocity. Therefore, the thrust of the engine must be equal to the force required to accelerate the air mass through the engine. Then, by using the symbol “Fn” for thrust pounds, the formula becomes: F = Ms (V2 - V1) G F = force in pounds Ms = mass flow in lb/sec V1 = inlet velocity V2 = jet velocity (exhaust) G = acceleration of gravity or 32.2 ft/sec2 where: change in velocity; difference between inlet velocity and jet velocity V2 - V1 = F = 100 lb/sec (800 ft/sec - 600 ft/sec) 32.2 ft/sec2 F = 621 lb F = 20,000 lb/sec 32.2 ft/sec2 Fn = Ms (V2 - V1) G Thrust of a gas turbine engine can be increased by two methods: increasing the mass flow of air through the engine or increasing the gas velocity. If the velocity of the turbojet engine remains constant with respect to the aircraft, the thrust decreases if the speed of the aircraft is increased. This is because V1 increases in value. This does not present a serious problem, however, because as the aircraft speed increases, more air enters the engine, and jet velocity increases. The resultant net thrust is almost constant with increased airspeed.

The Brayton cycle is the name given to the thermodynamic cycle of a gas turbine engine to produce thrust. This is a variable volume constant-pressure cycle of events and is commonly called the constant-pressure cycle. A more recent term is “continuous combustion cycle.” The four continuous and constant events are intake, compression, expansion (includes power), and exhaust. These cycles are discussed as they apply to a gas-turbine engine. In the intake cycle, air enters at ambient pressure and a constant volume. It leaves the intake at an increased pressure and a decrease in volume. At the compressor section, air is received from the intake at an increased pressure, slightly above ambient, and a slight decrease in volume. Air enters the compressor where it is compressed. It leaves the compressor with a large increase in pressure and decrease in volume, created by the mechanical action of the compressor. The next step, expansion, takes place in the combustion chamber by burning fuel, which expands the air by heating it. The pressure remains relatively constant, but a marked increase in volume takes place. The expanding gases move rearward through the turbine assembly and are converted from velocity energy to mechanical energy by the turbine. The exhaust section, which is a convergent duct, converts the expanding volume and decreasing pressure of the gases to a final high velocity. The force created inside the engine to keep this cycle continuous has an equal and opposite reaction (thrust) to move the aircraft forward.

Bernoulli’s principle (whenever a stream of any fluid has its velocity increased at a given point, the pressure of the stream at that point is less than the rest of the stream) is applied to gas turbine engines through the design of convergent and divergent air ducts. The convergent duct increases velocity and decreases pressure. The divergent duct decreases velocity and increases pressure. The convergent principle is usually used for the exhaust nozzle. The divergent principle is used in the compressor and diffuser where the air is slowing and pressurizing. 1-59 30 25 20 15 10 5 0 0 1 2 3 4 5 6 7 8 Compression ratio Thermal efficiency (in percent) Temperature 1600 °F1400 °F 1200 °F1000 °F 30 25 20 15 10 5 0 0 1 2 3 4 5 6 7 8 Compression ratio Thermal efficiency (in percent) Efficiency of C&T 75%85%80%90% efficiency.

Gas Turbine Engine Performance Thermal efficiency is a prime factor in gas turbine performance. It is the ratio of net work produced by the engine to the chemical energy supplied in the form of fuel. The three most important factors affecting the thermal efficiency are turbine inlet temperature, compression ratio, and the component efficiencies of the compressor and turbine. Other factors that affect thermal efficiency are compressor inlet temperature and combustion efficiency. Figure 1-81 shows the effect that changing compression ratio (compressor pressure ratio) has on thermal efficiency when compressor inlet temperature and the component efficiencies of the compressor and turbine remain constant. The effects that compressor and turbine component efficiencies have on thermal efficiency when turbine and compressor inlet temperatures remain constant are shown in temperature varies directly with turbine inlet temperature at a constant compression ratio.

Rpm is a direct measure of compression ratio; therefore, at constant rpm, maximum thermal efficiency can be obtained by maintaining the highest possible exhaust temperature. Since engine life is greatly reduced at high turbine inlet temperatures, the operator should not exceed the exhaust temperatures specified for continuous operation. Figure 1-83 illustrates the effect of turbine inlet temperature on turbine blade life. In the previous discussion, it was assumed that the state of the air at the inlet to the compressor remains constant. Since this is a practical application of a turbine engine, it becomes necessary to analyze the effect of varying inlet conditions on the thrust or power produced. The three principal variables that affect inlet conditions are the speed of the aircraft, the altitude of the aircraft, and the ambient temperature. To make the analysis simpler, the combination of these three variables can be represented by a single variable called stagnation density.

The power produced by a turbine engine is proportional to the stagnation density at the inlet. The next three illustrations show how changing the density by varying altitude, airspeed, and outside air temperature affects the power level of the engine. Figure 1-84 shows that the thrust output improves rapidly with a reduction in outside air temperature (OAT) at constant altitude, rpm, and airspeed. This increase occurs partly because the energy required per pound of airflow to drive the compressor varies directly with the temperature, leaving more energy to develop thrust. In addition, the thrust output increases since the air at reduced temperature has an increased density. The increase in density causes the mass flow through the engine to increase. The altitude effect on thrust, as shown in Figure 1-85, can also be discussed as a density and temperature effect. In this case, an increase in altitude causes a decrease in pressure and temperature.

Since the temperature lapse rate is lower than the pressure lapse rate as altitude is increased, the density is decreased. Although the decreased temperature increases thrust, the effect of decreased density more than offsets the effect of the colder temperature. The net result of increased altitude is a reduction in the thrust output. The effect of airspeed on the thrust of a gas-turbine engine is shown in Figure 1-86. To explain the airspeed effect, it is necessary to understand first the effect of airspeed on the factors that combine to produce net thrust: specific thrust and engine airflow. Specific thrust is the net thrust in pounds developed per pound of airflow per second. It is the remainder of specific gross thrust minus specific ram drag. As airspeed is increased, ram drag increases rapidly. The exhaust velocity remains relatively constant; thus, the effect of the increase in 1-60 0 100 200 300 400 500 Turbine bucket life (in hours) Turbine inlet temperature Increase Dec STD INC Outside air temperature (OAT) Net thrust 60% 100% Constant altitude, rpm, and airspeed life.

0 10 20 30 40 50 Altitude—thousand of ft Thrust—percent sea level rated 100 50 0 Constant airspeed and rpm 0 TAS Net thrust Increase 100 50 0 Constant rpm, altitude, and OAT With ram Without ram airspeed results in decreased specific thrust. [Figure 1-86] In the low-speed range, the specific thrust decreases faster than the airflow increases and causes a decrease in net thrust. As the airspeed increases into the higher range, the airflow increases faster than the specific thrust decreases and causes the net thrust to increase until sonic velocity is reached. The effect of the combination on net thrust is illustrated in Figure 1-87.

Ram Recovery A rise in pressure above existing outside atmospheric pressure at the engine inlet, as a result of the forward velocity of an aircraft, is referred to as ram pressure. Since any ram effect causes an increase in compressor entrance pressure over atmospheric, the resulting pressure rise causes an increase in the mass airflow and gas velocity, both of which tend to increase thrust. Although ram effect increases engine thrust, the thrust being produced by the engine decreases for a given throttle setting as the aircraft gains airspeed. Therefore, two opposing trends occur when an aircraft’s speed is increased.

What actually takes place is the net result of these two different effects. An engine’s thrust output temporarily decreases as aircraft speed increases from static, but soon ceases to decrease. Moving toward higher speeds, thrust output begins to increase again due to the increased pressure of ram recovery. 1-61 0 TAS Increase Airflow (in lb/sec) Specific thrust Total airflow (w a) Constant rpm, altitude, and OAT airflow.