Temperature and pressure notation of a typical turbo-jet engine

INTRODUCTION

Fig. 21-1 Temperature and pressure notation of a typical turbo-jet engine.

1. The performance requirements of an engine are obviously dictated to a large extent by the type of operation for which the engine is designed. The power of the turbo-jet engine is measured in thrust, produced at the propelling nozzle or nozzles, and that of the turbo-propeller engine is measured in shaft horse-power (s.h.p.) produced at the propeller shaft. However, both types are in the main assessed on the amount of thrust or s.h.p. they develop for a given weight, fuel consumption and frontal area.

2. Since the thrust or s.h.p. developed is dependent on the mass of air entering the engine and the accel- eration imparted to it during the engine cycle, it is obviously influenced, as subsequently described, by such variables as the forward speed of the aircraft, altitude and climatic conditions, These variables influence the efficiency of the air intake, the compressor, the turbine and the jet pipe; conse- quently, the gas energy available for the production of thrust or s.h.p. also varies.
3. In the interest of fuel economy and aircraft range, the ratio of fuel consumption to thrust or s.h.p. should be as low as possible. This ratio, known as the specific fuel consumption (s.f.c.), is expressed in pounds of fuel per hour per pound of net thrust or s.h.p. and is determined by the thermal and propulsive efficiency of the engine. In recent years considerable progress has been made in reducing s.f.c. and weight. These factors are further explained in para. 46.
4. Whereas the thermal efficiency is often referred to as the internal efficiency of the engine, the propulsive efficiency is referred to as the external efficiency. This latter efficiency, described in para. 37,explains why the pure jet engine is less efficient than  the turbo-propeller engine at lower aircraft speedsl eading to development of the by-pass principle and, more recently, the propfan designs.
5. The thermal and the propulsive efficiency also influence, to a large extent, the size of the compressor and turbine, thus determining the weight and diameter of the engine for a given output.
6. These and other factors are presented in curves and graphs, calculated from the basic gas laws (Part 2), and are proved in practice by bench and flight testing, or by simulating flight conditions in a high altitude test cell. To make these calculations, specific symbols are used to denote the pressures and tem- peratures at various locations through the engine; for instance, using the symbols shown in fig. 21-1 the overall compressor pressure ratio is   . These  symbols vary slightly for different types of engine; for instance, with high by-pass ratio engines, and also  when afterburning (Part 16) is incorporated,additional symbols are used.
7. To enable the performance of similar engines to be compared, it is necessary to standardize in some conventional form the variations of air temperature and pressure that occur with altitude and climatic conditions. There are in use several different definitions of standard atmospheres, the one in most common use being the International Standard Atmosphere (I.S.A.). This is based on a temperature lapse rate of approximately 1.98 K. degrees per 1,000ft,, resulting in a fall from 288.15 deg.K. (15 deg.C) at sea level to 216.65 deg.K (-56.5 deg.C.) at 36,089 ft. (the tropopause). Above this altitude the 1 3 P P temperature is constant up to 65.617ft. The I.S.A. standard pressure at sea level is 14.69 pounds per  square inch falling to 3.28 pounds per square inch atthe tropopause (refer to I.S.A. table fig. 21-10).