For the past 50 years, the principal propulsion source for military and civil aircraft has been the gas turbine. For a variety of technical reasons, this situation is likely to continue into the foreseeable future. The earliest military aircraft gas turbines, developed toward the end of World War II, opened the way to high-speed flight by providing high power from low weight, compact engines. For military use, these "jet" engines offered an escape from the aerodynamic limitation of the propeller, despite their relatively low thermodynamic efficiency at the time. Early civil aircraft gas turbines continued, at first, to use the engine to drive a propeller in so-called turbo-prop form (engine shown in Figure 7-7a). Several types of turbo-props are still used for short-haul operations where cruise speeds are less important, but larger and faster aircraft have dispensed with the propeller.
Since those early days, huge strides have been made in the critical and pacing technology fields that influence the key design and performance characteristics of engine design. In particular, major advances have been made in the fields of turbo-machinery aerodynamics, combustion, turbine blade cooling, and materials. For military engines, these advances have been realized mainly in increases in the ratio of engine thrust to engine weight. For civil aircraft engines, the benefits have led to high bypass ratio engines with substantially lower fuel consumption, which have contributed to the rapid growth in air transportation over the past 3 decades.
It is common to compare engines in terms of specific fuel consumption (SFC), which is the fuel flow rate per unit thrust at cruise. However, the ultimate goal is to minimize total fuel burned per unit payload, rather than SFC; this computation involves engine weight, installation drag, and their effect on the total fuel required to complete a flight mission. For subsonic transport aircraft, the weight of the engines is on the order of 10-15% of the empty weight of the aircraft; a reduction of one unit of total engine weight translates to a reduction of between 1.5 and 4 units of aircraft empty weight, depending on the design. The relatively larger reduction in aircraft weight derives from concomitant reductions in requirements for supporting structure. The benefits are further magnified by the fact that the reduction in fuel burn attributable to engine weight savings is proportional to increasing aircraft range.
The following subsections present a brief outline of engine performance issues from an historical perspective (7.4.2) and a look forward into the future (7.4.3). The link between performance considerations and emissions is discussed in Section 7.4.4. Because the gas turbine is expected to remain the principal power source for aircraft propulsion well into the future, however, Section 7.4.1 presents a simplified review of the engine's fundamental principles to explain, among other things, why there is no reasonable alternative to the gas turbine or derivatives thereof in the foreseeable future.
Figure 7-8: Gas turbine thermal efficiency.
The core of the basic gas turbine consists of three essential elements: The compressor, which mechanically increases the energy of the air (raising the pressure and temperature); the combustor, in which fuel is burned (further raising the temperature of the pressurized air); and the turbine, which mechanically extracts enough energy from the hot compressed gas to drive the compressor (thereby reducing the pressure and temperature of the gas). A fraction of the net energy remaining in the gas after it leaves the turbine is then available to be used in different ways, as shown in Figure 7-7. Case (a) uses an additional turbine stage to mechanically convert the energy to shaft work (such as might be required to drive a propeller or electric generator); case (b) is the turbojet, in which a nozzle is used to accelerate the gas (converting some of the energy into kinetic energy), producing a high-speed jet that can be used to propel the vehicle; and case (c) is the turbofan, in which a further turbine converts most of the energy of the gas into shaft work to drive the bypass air compressor, thereby producing the bypass jet that propels the vehicle. For aircraft applications, the weight of the engine must be low in relation to the power output. This constraint has kept the aircraft type of gas turbine simple-much simpler than those now being built for land-based power generation. (The implications of going to more complicated configurations are considered briefly in Section 7.4.4.)The overall efficiency of an aircraft engine, h0, is the mechanical power created by the thrust divided by the energy input rate of the fuel flow. It is convenient to express overall efficiencies by h0 = htherm x hp, where htherm is the thermal efficiency and hp is the propulsion efficiency; these terms are considered below.
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