DESCRIPTION OF THE GAS TURBINE ENGINE

Basically, the gas turbine engine consists of a compressor, a combustion chamber, and a turbine. The combination of these basic components is referred to as the gas generator or core engine. Other components are then added to make the complete engine. It is beyond the scope of this text to delve into the details of gas turbine engine design. However, the various types of gas turbine engines will be described, and their operating characteristics will be discussed in some detail.

Beginning with the core engine, the turbojet engine pictured in Figure 6.23 is obtained by adding an engine air inlet and a jet nozzle. As the air

enters the inlet, it is diffused and compressed slightly. It then passes through a number of blade rows that are alternately rotating and stationary. The collection of rotating blades is referred to as the rotor; the assembly of stationary blades is called the stator. This particular compressor configuration is known as an axial-flow compressor and is the type used on all of today’s larger gas turbine engines. Early gas turbine engines, such as Whittle’s engine, employed a centrifugal compressor, as shown in Figure 6.24. Here, the air enters a rotating blade row near the center and is turned radially outward. As the air flows out through the rotating blade passage, it acquires a tangential velocity component and is compressed. A scroll or radial diffuser collects the compressed air and delivers it to the combustion chamber. Centrifugal com­pressors were used on the early turbojet engines simply because their design was better understood at the time. As jet engine development progressed, centrifugal compressors were abandoned in favor of the more efficient axial – flow compressors. The axial-flow compressor also presents a smaller frontal area than its centrifugal counterpart and is capable of achieving a higher pressure ratio.

Smaller sizes of gas turbine engines still favor the centrifugal com­pressor. Figure 6.25 is a cutaway drawing of the Garrett ТРЕ 331/T76 turboprop engine. The compressor section of this engine consists of two stages of radial impellers made of forged titanium.

•(a) <b)

Figure 6.24 Typical centrifugal flow compressor impellers, (a) Single-entry im­peller. (b) Double-entry impeller. (Courtesy General Electric Co.)

Figure 6.25 Cutaway of Garrett TPE 331/76. (Courtesy The Garrett Corp.)

After the compressed air leaves the compressor section, it enters the combustor, or burner, section. Atomized fuel is sprayed through fuel nozzles and the resulting air-fuel mixture is burned. Typically, the ratio of air to fuel by weight is about 60:1. However, only approximately 25% of the air is used to support combustion. The remainder bypasses the fuel nozzles and mixes downstream of the burner to cool the hot gases before they enter the turbine.

The mixed air, still very hot (about 1100 °С), expands through the turbine stages, which are, composed of rotating and stationary blade rows. The turbines extract energy from the moving gases, thereby furnishing the power required to drive the compressor. Nearly 75% of the combustion energy is required to drive the compressor. The remaining 25% represents the kinetic energy of the exhaust, which provides the thrust. For example, in the General Electric CF6-6 turbofan engine [180,000 N (40,0001b) thrust class], the turbine develops approximately 65,600 kW (88,000 shp) to drive the high – and low – pressure compressors.

Variations of the gas turbine engine are presented in Figure 6.26. In a turboprop or turboshaft engine, nearly all of the energy of the hot gases is extracted by the turbines, leaving only a small residual thrust. The extracted energy in excess of that required to drive the compressor is then used to provide shaft power to turn the propeller or a power-output shaft in general. Turboshaft engines power most of today’s helicopters and are used exten­sively by the electric utilities to satisfy peak power load demands.

A “spool” refers to one or more compressor and turbine stages con­nected to the same shaft and thus rotating at the same speed. Gas turbine engines generally use one or two spools and are referred to as single or dual compressor engines. A turboshaft engine may incorporate a free turbine that is independent of any compressor stage and is used solely to drive the shaft. Since the rotational speed of a turbine wheel is of the order of 10,000 rpm, a
reduction gear is required between the turbine shaft and the power output shaft.

A turboprop produces a small amount of jet thrust in addition to the shaft power that it develops; these engines are rated statically in terms of an equivalent shaft horsepower (eshp). This rating is obtained by assuming that 1 shp produces 2.5 lb of thrust. For example, the dash 11 model of the engine shown in Figure 6.25 has ratings of 1000 shp and 1045 eshp. From the definition of eshp, this engine therefore produces a static thrust from the turbine exhaust of approximately 1131b.

A turbojet engine equipped with an afterburner is pictured in Figure 6.26/. Since only 25% or so of the air is used to support combustion in the burner section, there is sufficient oxygen in the turbine exhaust to support additional burning in the afterburner. Both turbofans and turbojets can be equipped with afterburners to provide additional thrust for a limited period of time. Afterburning can more than double the thrust of a gas turbine engine, but at a proportionately greater increase in fuel consumption. Essentially, an afterburner is simply a huge stovepipe attached to the rear of an engine in lieu of a tail pipe and jet nozzle. Fuel is injected through a fuel nozzle arrange­ment called spray bars into the forward section of the afterburner and is

ignited. This additional heat further expands the exhaust, providing an in­creased exhaust velocity and, thereby, an increased thrust. The afterburner is equipped with flame holders downstream of the spray bars to prevent the flames from being blown out of the tail pipe. A flame holder consists of a blunt shape that provides a wake having a velocity that is less than the velocity for flame propagation. An adjustable nozzle is provided at the exit of the afterburner in order to match the exit area to the engine’ operating condition.

Two different types of turbofan engines are shown in Figure 6.26b and 6.26c; the forward fan with a short duct and the forward fan with a long duct. These engines are referred to as bypass engines, since part of the air entering the engine bypasses the gas generator to go through the fan. The ratio by weight of the air that passes through the fan (secondary flow) to the air that passes through the gas generator (primary flow) is called the bypass ratio. Early turbofan engines had bypass ratios of around 1:1; the latest engines have ratios of about 5:1. One such engine, Pratt & Whitney’s JT9D turbofan, is shown in Figure 6.27. Included on the figure are temperatures and absolute pressures throughout the engine for static operation at standard sea level conditions.