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Aviation History
1977
1977 - 0125.PDF
FLIGHT International, 15 January 1977 115 cent of the material purchased is removed during manu facture of engine parts. Such a figure suggests that there is great scope for cost reductions through the use of alternative manufacturing processes. Development should lead to the increased use of such processes as precision casting, extrusion and rolling in preference to machining. New welding developments such as inertia bonding and laser-beam welding can also be expected to contribute to future cost savings. Exhaust-emission regulations for civil aero-engines have now been framed. While emissions of carbon monoxide, unburnt hydrocarbons and unburnt carbon can be reduced through attention to the combustion process, the oxides of nitrogen present greater difficulties. The research effort now directed at this problem includes investigations of lower-temperature combustion systems producing the minimum of oxides of nitrogen. Up to now, the control systems of most aero gas-turbines have been hydro-mechanical. Much interest now centres on digital electronics, which offer highly sophisticated control, great control-law flexibility (by the use of suitable soft ware), some hardware commonality between different types of engine, and compatibility with other digital equip ment such as air-data or flight-control computers. Digital systems are now beginning to find their way into aircraft powerplants—the Concorde variable air intake, for example. It seems likely that digital control will in time supersede other forms wherever the control requirements are complex. Although the modern civil turbof an is much quieter than earlier jet engines, our understanding of several aspects of engine noise is still incomplete. As a result, there is some uncertainty in predicting the noise from a projected engine, particularly if it differs substantially from its predecessors. Continuing research can be expected to improve this situation. In the field of instrumentation, the further development of telemetry from rotating components, X-ray observation of seal movements during engine transients, noise-source location by advanced microphone and signal-processing systems, laser anemometry for turbo-machine flows, and several other advanced measuring and observing tech niques will yield imporant benefits. Propulsive efficiency is improved by reducing the velocity of the jet in relation to the aircraft speed—i.e. by reducing the thrust per unit airflow, or "specific thrust," of the powerplant. Thermal efficiency, on the other hand, is related to the thermodynamics of the engine and depends on the cycle and the individual component efficiencies. For the basic gas-generator section of the engine, thermal efficiency is improved by increasing pressure ratio, turbine entry temperature and component efficiency. If a bypass system is used, the losses associated with transferring energy from the gas generator to the bypass stream have an adverse effect on the thermal efficiency of the engine as a whole. When overall efficiency is plotted in relation to thermal efficiency, propulsive efficiency and the ratio specific thrust/flight speed, these points emerge: • Very substantial improvements in both thermal and propulsive efficiency in subsonic flight have been achieved since the inception of jet propulsion. • Although the Olympus 593 and RB.211 differ radically in design, both show a propulsive efficiency of about 75 per cent at their respective cruise conditions—i.e. the ratio jet velocity/flight velocity is approximately the same in both cases. This serves to illustrate the suitability of the simple turbojet for high-speed flight, in which its high jet velocity is well matched to the conditions. • Despite the much higher pressure ratio of the three- shaft RB.211 compared with the two-shaft Olympus (27 as compared with 11), the latter shows a markedly superior thermal efficiency of no less than 55 per cent. This stems from the augmentation of the basic Olympus engine cycle by ram compression due to forward speed, giving an effective cycle pressure ratio which exceeds 80 at Mach 2. The above comparisons give some impression of the relative difficulty of achieving high efficiency with jet propulsion at subsonic flight speeds. A low specific-thrust 30- O CIVIL • MILITARY 25 PRESS. RATIO 20r i 15 IO 5 SPE^5I2QQ||;TF34 **TF4I-AI AVON 350 ^ifllwfr LARZAC 04 J 85 W M 53 GYRON JUNIOR^ , An GYRON-"- U fcffp;0RPH£lJS3 w 8-11 STAGES TOTAL 1945 1955 1965 FIRST FLIGHT 1975 Figure 3: the progress of pressure ratio. Figure 4 (below): the progress of turbine-entry temperature TURBINE ENTRY TEMPERATURE ~°K 1800 1600 WOO 1200 1000 . - • A' A Mr l$*r .' IP^NCOOLED COOLED 1945 1955 1965 1975 YEAR OF ENTRY INTO SERVICE 1985 value is needed to get high propulsive efficiency, calling for the transfer of a substantial proportion of the gas- generator power to a bypass stream. And on the thermal efficiency side, the augmentation of cycle pressure ratio by ram compression is quite limited. A high compression capability therefore has to be built into the engine itself. Compared with the early jets, modern subsonic engines have gone some way in both these directions, with bypass ratios of about 5 and cycle pressure ratios now around 30. In this development, design complexity has been kept within acceptable limits by advances in turbo-machinery aerodynamics, materials and mechanical design. This is evidenced in figure 3, which shows the progression of cycle pressure ratio for axial-flow engines. Initially the number of compressor and turbine stages was increased, but that process did not go beyond a total of about 20 stages. Subsequently, more and more "performance" in terms of pressure ratio has been wrung out of this number of stages, while component efficiency levels have been main-
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