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Aviation History
1956
1956 - 0562.PDF
562 FLIGHT THE SUPERSONIC TURBOJET . . . it is essential that the balance of the flow around the peripheryof the compressor should not be disrupted. Air for cooling the turbine can probably best be extracted fromthe wall of the gas path—the inner wall seems logical from the geometrical viewpoint—and it should be possible to peel off therequired flow without greatly disturbing the flow pattern. It does at least seem to be a case in which a design, having been gotto work, stays correct. General increases in gas temperature so far do not seem to have caused much difficulty in cooling theprimary zone walls. Even at over 2,000 deg K it is quite possible 10 maintain satisfactory metal temperatures by film cooling withsecondary or tertiary air which can form a thermal barrier be- tween the hottest gas and the turbine nozzle box. As already described at some length, turbine-inlet tempera-tures must be allowed to rise in order to obtain the maximum specific output. In the fairly-high-pressure VG type of enginetwo turbine stages may be necessary, and both would have to be cooled in high-speed flight. Engineers on both sides of theAtlantic have now evolved several forms of turbine-disc and blade assembly which lend themselves well to air cooling. One THE VG ENGINE This hypothetical design of turbojet is extensively discussed in the text. It is intended to meet the requirements of efficient cruise at about Mach 0.9 while providing for bursts of speed up to Mach 2.5-2.7. Insets show: A, variable spike; B, cowl flaps shut; C, cowl moved forward; D, nozzle moved to fully expanded position; and E, aerodynamic constriction of the throat. 1 Extendable noie (see sketch A). 2 Water/methanol rotary spray (anti-icing/intake cooling). 3 Nose portion slides forward from this point, increasing the primary intake area and pro- viding peripheral bleed (see sketch C). 4 Shutters move forward and contract, sealing intake (sketch B). 5 Nose-cone ram. e Pumps and accessories. 7 Electronics. 9 Rams and rails for sliding nose section. f Integral oil tank. 10 Casing/cowling splits into main units for servicing. 11 Air trunks. 12 Exit guide vanes and centre- bearing housing. 13 Air bleeds. 14 Air-bleed control valves. 15 Air trunk to 17 and 18. 1* Cooled blading. 17 Gate valve: air supply to throat- varying nozzles. 18 Gate valve controlling air- driven motors (20). 19 Sliding portion for increased nozzle area (see sketch D). 20 Air-motor screw-jacks for slid- ing section (19). 21 Air manifold. 22 Jets for aerodynamic throat- area control (sketch E). 23 Afterburner step-type flame- holder. 24 Afterburner fuel-spray nozzles. afterburners would be operated "nearly dry" with an almost 2:1gain in six. compared with present afterburning engines. One point which cannot be overemphasized is that, as the flightMach number is increased beyond about 1.8, with an afterburner in operation the cycle efficiency of the basic engine becomes pro-gressively less important. The thrust and consumption depend overwhelmingly on the performance of the afterburner, and itbecomes theoretically quite practicable to pull out half the blades from the engine without spoiling the performance; in fact, itshould even improve it. Nevertheless, a major design-postulate of the hypothetical VG engine is that it should be suitable forsubsonic cruise; hence the annoying requirement of high, built- in pressure ratio. In general the supersonic afterburner will have to meet thesame sort of requirements as those facing present units, although truly formidable throughput and flow velocity will accentuatemost problems, particularly that of minimizing pressure drop. At the Lewis Flight Propulsion Laboratory the N.A.C.A. havedevoted much research-effort to the evolution of new afterburner fuels in which the speed of the flame front is much greater than of these forms embodies twin discs for each stage; not only cancooling-air be readily fed between the discs, but such a rotor may well have better mechanical properties than single-disc assembliesof the same weight. Naturally such a design would require a com- pletely new form of blade root. Two basic factors which determine the flow of cooling airthrough the blades are the pressure differences across the blading and the total resistance to the flow. At high altitudes and tempera-tures it may not be an easy matter to maintain a sufficiently high differential to keep the blade temperatures down unless assistances provided by the centrifugal acceleration (up to 20,000 g) within he turbine rotor blades themselves. The latter effect, of course,s absent from the nozzle guide vanes or stator blades—where, urthermore, the discharge has to be made into a relatively highngnation pressure. The problem can be partially eased by eject- ng the cooling air from the blade trailing edges, insofar as thiss mechanically possible; in the rotor the much greater pressure differential puts the designer in a better position even discount-ng the centrifugal assistance. In order to prevent the overall diameter from becoming exces-;ive, supersonic turbojets are likely to operate with an increased low Mach number between the turbine and the propelling nozzle.Nevertheless, the designer may be in trouble in the exit cone at he upstream end of this section, and the diffusion rate may haveo be reduced if losses are to be kept acceptable. This inevitably ncreases the length of the engine. As already noted, the VG design is a fairly high-pressurefterburning unit intended to compromise between supersonic and ubsonic requirements. Unfortunately, at high supersonic flighttlach numbers the increased temperature of the flow through- >ut the engine reduces the margin available for the combustion•f fuel behind the turbine, so that the maximum available after- lurning boost is proportionately lower. A General Electricngineer, in making this point last summer, said that future in the hydrocarbon fuels at present employed. By employing suchfuels (in the afterburner only) reheat can be accomplished with very much reduced drag from flameholders and other baffleassemblies. Much depends, of course, on how much afterburning is neces-sary to meet the design requirements. Every gallon of fuel con- sumed in the afterburner would theoretically do more good if itwere burned upstream of the turbine, and that is certainly where all the fuel would be put if it were not for the old bogey of theceiling on turbine-inlet temperature. With a fairly high-pressure engine of the VG type, the ideal aircraft installation would accom-modate an afterburner of massive proportions—perhaps 6ft in diameter and 30ft long—in which the mean gas temperature couldbe reheated to more than 2,000 deg K with minimum losses. Few aircraft seem likely to provide this sort of space-envelope, and,where the afterburner is dimensionally limited, the percentage boost will have to be correspondingly cut down in order to reducethe pipe Mach number and consequent losses. In the classic subsonic engine, of the type evolved by Whittleand now used throughout the world, acceptable propulsion efficiency is obtained merely by letting the jet escape through asimple outlet terminating at a wholly subsonic propelling nozzle. The latter may be tailored to the correct area giving optimumthrust for acceptable jet-pipe temperature by having local restric- tors (colloquially termed "mice") welded around the periphery.In contrast, the supersonic turbojet will employ a more sophisti- cated type of nozzle, of the convergent-divergent (or con-di)form which is already familiar from existing designs of rocket motor. In such a nozzle the flow is restricted by a throat and isthen expanded in a supersonic divergent portion so that, as the pressure returns to that of the surrounding atmosphere, the jetis accelerated to a very high speed indeed. Con-di nozzles of any given form reach a position of equi-librium at either of two Mach numbers; for a typical nozzle the
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