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Aviation History
1950
1950 - 1384.PDF
IO4 FLIGHT, 20 July 1950 POWER FOR THE FUTURE Relative Merits of Pbton Engines, Turboprops and Turbojets Reviewed at World Power Conference ONE of the more important symposia in the inter-national calendar came to a conclusion last week inLondon. This, the fourth in the series of world conferences on power—the first (1924) was also held in London, the second (1930) was in Berlin, and the third (1936) was in Washington—was of considerable im- portance, not only by virtue of the common meeting-ground it afforded to technicians of many nationalities, but also because of the large number of notable contributions to the programme of learned papers presented. We publish bereunder a digest of the paper, Power Plant Requirements for Future Aircraft, by Messrs. Owner and Hooker,* given in the Great Hall of the Institute of Civil Engineers on July 14th. In the introduction to their paper, the authors stated that the classic simplification made by Sir Frank Whittle in using the same air in the thermodynamic cycle and the propulsive mechanism had led to no small confusion of thought in the assessment of the effects which changes in design parameters had on efficiency. A notable example of this could be found in the prevalent impression that raising the combustion tem- perature improved the efficiency of turbojets, whereas, of course, although the thermal efficiency of the cycle was improved by an increase in temperature, the propulsive efficiency was reduced by the higher jet velocity, and the combination gave a net reduction in overall efficiency. Specific power output was, of course, another thing entirely. For the purpose of subsequent discussion, definitions of efficiency were given: Cycle thermal efficiency was the per- centage of the fuel power which was made available for con- version to net propulsive power by the main thermal Cycle. Propulsive efficiency was the percentage of the gas horse- power produced by the tbermodynamic cycle, which was converted into net propulsive thrust power acting on the aircraft. Overall efficiency was_ the percentage of fuel power which appeared as net propulsive power. (It could be noted that, by definition, the overall thermal efficiency was the product of the propulsive and cycle thermal efficiencies as defined.) The power plant could be regarded as being divided into a main thermal cycle for the production of power in the form of heated and compressed air, and into a propulsive mechanism for converting this power into thrust acting upon the aircraft. Given the component efficiencies of the compressor, the com- *F. M. Owner, C.B.E.. M.Sc, F.R.Ae.S., M.S.A.E., and S. G. Hooker, O.B.E., D.Phil., B.Sc, F.R.Ae.S.. are respectively chief engineer and assistant chief engineer of the Engine Division, the Bristol Aeroplane Co., Ltd. bustion system and the turbine for driving the compressor, the only remaining parameters upon which the thermal efficiency of the cycle depended were (i) the compression ratio and (ii) the peak cycle temperature. As a result of the necessity to keep the bulk and weight of units to a minimum, the efficiency of compressors of practical design fell with increasing compression ratio at a rate greater than would be anticipated from a constant stage polytropic efficiency. The assumed fall in compressor efficiency shown in Fig. 1 represented efficiency values that could be obtained by employing the best modern technique of axial compressor design. The ability to start the engine easily, to accelerate it and to operate at part load without surging (all essential requirements for aircraft power plants) necessitated com- pounding axial compressors for compression ratios appreciably in excess of 6:1. The variation of cycle thermal efficiency with compression ratio, forward speed and altitude was shown in Fig. 1, and it was significant that, under all conditions, the optimum thermal efficiency was obtained at compression ratios between 9 and 10:1. In view of the aerodynamic difficulties in high compres- sion-ratio compressors, it was naturally advisable to adopt the lowest permissible compression ratio and, consequently, the authors had taken a compression ratio of 8.5:1 for their calcu- lations as representing a good practical compromise. The main reason for the increase in thermal efficiency with forward speed and altitude evident in Fig. 1 was the increase in compression ratio due to the ram compression in the air intake, which had been computed at 90 per cent, isentropic efficiency. The influence of altitude was secondary, because it only reduced the temperature of the air entering the intake. For example, at 1,000 m.p.h. at 40,000ft, the fall in intake temperature due to altitude would be 72 deg C, whilst the rise in intake temperature due to forward speed would be 100 deg C ••> Consequently, the net air-intake temperature was increased by 28 deg C over the static sea level value; despite this, how- ever, the thermal efficiency was increased from 32 to 48 per cent. The design of the air intake was therefore clearly a matter of vital importance. The remaining parameter affecting the main cycle thermal efficiency was the peak' cycle temperature, and its effect was shown in Fig. 2, the curve having been computed for a designed compression ration of 8.5:1. It was not as yet pos- sible to exceed a peak cycle temperature of more than 1,150 deg. K for protracted periods, owing to turbine-blade limita- tions. It was, however, clear from the diagram that the gain in thermal efficiency would be progressive and, further, the. increase in specific output for any given power plant would increase enormously the peak cycle temperature—a point of GAS TEMPERATURE HOO*K TURONE EFFK3ENC* «7 B » £S 8 14 sea -tern. p«ssu«e awio Fig. I left). Curves of cycle thermal efficiency for ras turbines. rig. J, zmye . uycie t&srniai efficiency at various peak cycte temperatures.
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