438
12 Thrust Gas Turbines
Thrust specific fuel consumption thus depends on the flight speed. Values must
be compared with the same flight Mach number. This is not a problem in practice.
With airliners, the flight Mach number is always about 0.85. With fighter aircraft,
the design flight Mach number is typically around 2. But, there are also applications
with other flight Mach numbers. An example is paratrooper transport aircraft with a
typical flight Mach number around 0.5. Note that thrust specific fuel consumption
has meaning in static conditions (start of take-off at sea level) where efficiency has
no meaning.
12.4 Performance of the Gas Generator
and the Single-Jet Engine
The gas generator in an aero gas turbine is analogous with a simple-cycle power gas
turbine. Good performance therefore requires a high pressure ratio and a high temperature ratio. High pressure ratio is important for thermodynamic efficiency. The
temperature ratio mainly determines the specific power. Compared to power gas
turbines, aero gas turbines show an important advantage concerning pressure ratio
and temperature ratio. Compression is partially obtained aerodynamically by the
ram effect of the incoming air within the engine inlet. The ram effect is the spontaneous deceleration due to the obstruction formed by the engine. With an airliner, the
aerodynamic compression achieves a very high efficiency. Further, it does not require any corresponding turbine work. Of course, the ram effect causes drag for the
engine and for the aircraft. The through-flow Mach number (Mach number of the
average meridional velocity component) in an aero-engine compressor is typically
around 0.5. In a fan, it is typically higher, around 0.6. The flight Mach number of
an airliner is 0.8–0.85. So, the deceleration at the engine entrance is very modest.
Hence, the isentropic efficiency of the inlet is very high, of the order of 0.997. For
analysis purposes, we will assume it equal to unity. For a supersonic fighter aircraft,
the efficiency of the inlet is much lower due to formation of shock waves. The atmosphere temperature is low at cruising altitude: 225 K at 10 km height. At a given
combustion temperature, the temperature ratio is thus better than with a power gas
turbine. Moreover, the final temperature after compression is lower due to the lower
atmospheric temperature, resulting in a more efficient turbine cooling. A turbojet
engine consists of a gas generator with a directly joined nozzle. Nozzle efficiency
amounts to about 0.98, being very close to 1. The gas generator may be described
as a single-jet engine with an ideal nozzle. Gas power and dynamic power are then
identical. Henceforth we analyse the single-jet engine with an ideal nozzle. The
results strictly apply to the gas generator. Thermal efficiency and specific power
values are about 2 % lower with a real turbojet engine. As with a power gas turbine,
the major loss in a turbojet engine is thermal, in the sense of remaining heat in the
combustion gas after expansion in the nozzle. In second order, there are thermodynamic losses due to irreversibility in the compressor and the turbine and due to
cooling of turbine parts. As with the power gas turbine it is therefore relevant to
accomplish a first analysis with neglect of thermodynamic losses.
12 Thrust Gas Turbines
Thrust specific fuel consumption thus depends on the flight speed. Values must
be compared with the same flight Mach number. This is not a problem in practice.
With airliners, the flight Mach number is always about 0.85. With fighter aircraft,
the design flight Mach number is typically around 2. But, there are also applications
with other flight Mach numbers. An example is paratrooper transport aircraft with a
typical flight Mach number around 0.5. Note that thrust specific fuel consumption
has meaning in static conditions (start of take-off at sea level) where efficiency has
no meaning.
12.4 Performance of the Gas Generator
and the Single-Jet Engine
The gas generator in an aero gas turbine is analogous with a simple-cycle power gas
turbine. Good performance therefore requires a high pressure ratio and a high temperature ratio. High pressure ratio is important for thermodynamic efficiency. The
temperature ratio mainly determines the specific power. Compared to power gas
turbines, aero gas turbines show an important advantage concerning pressure ratio
and temperature ratio. Compression is partially obtained aerodynamically by the
ram effect of the incoming air within the engine inlet. The ram effect is the spontaneous deceleration due to the obstruction formed by the engine. With an airliner, the
aerodynamic compression achieves a very high efficiency. Further, it does not require any corresponding turbine work. Of course, the ram effect causes drag for the
engine and for the aircraft. The through-flow Mach number (Mach number of the
average meridional velocity component) in an aero-engine compressor is typically
around 0.5. In a fan, it is typically higher, around 0.6. The flight Mach number of
an airliner is 0.8–0.85. So, the deceleration at the engine entrance is very modest.
Hence, the isentropic efficiency of the inlet is very high, of the order of 0.997. For
analysis purposes, we will assume it equal to unity. For a supersonic fighter aircraft,
the efficiency of the inlet is much lower due to formation of shock waves. The atmosphere temperature is low at cruising altitude: 225 K at 10 km height. At a given
combustion temperature, the temperature ratio is thus better than with a power gas
turbine. Moreover, the final temperature after compression is lower due to the lower
atmospheric temperature, resulting in a more efficient turbine cooling. A turbojet
engine consists of a gas generator with a directly joined nozzle. Nozzle efficiency
amounts to about 0.98, being very close to 1. The gas generator may be described
as a single-jet engine with an ideal nozzle. Gas power and dynamic power are then
identical. Henceforth we analyse the single-jet engine with an ideal nozzle. The
results strictly apply to the gas generator. Thermal efficiency and specific power
values are about 2 % lower with a real turbojet engine. As with a power gas turbine,
the major loss in a turbojet engine is thermal, in the sense of remaining heat in the
combustion gas after expansion in the nozzle. In second order, there are thermodynamic losses due to irreversibility in the compressor and the turbine and due to
cooling of turbine parts. As with the power gas turbine it is therefore relevant to
accomplish a first analysis with neglect of thermodynamic losses.