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Aviation History
1955
1955 - 1577.PDF
694 FLIGHT, 28 October 1955 700 I X / A / PISTOM EN( / / iMES / jn am .LAUD aSUfl INE3. fLAN£S AMES KM* Mil IWO 1920 I93O 1940 IS50 IMO 1370 THRUST AT SOJOO (OF ENGINE GJVIMG taOOOL AT SEA LEVEL STATIC. )„ WOO ZOOO J00 400O fooo *t>uo 7000 STILL AIR LENGTH/ N. MILE FLIGHT MACti NUMbEft. The progress of World Speed Records, Fig. 2 (far left) includes the 822 m.p.h. (F-100C) figure now awaiting ratifica- tion. Fig. 3 is a plot of intake efficiencies and Fig. 4 (above) compares transport engine direct operating costs. PROPULSION PROGRESS . . . . of 93 per cent. The figure also shows that the pure jet, with anefficient intake, can achieve quite a high overall efficiency at high Mach number, e.g., 40 per cent at about M = 2.5. To maintain theoverall efficiency at high flight speeds it is essential to operate at high flame temperatures. We have seen that this is dependent on someform of blade cooling, which is made difficult by the high temperature of the cooling air available. "A further development which should pay handsome dividends athigh speeds is the convergent-divergent nozzle. At the moment, most engines operate with a convergent nozzle in which the final jet velocityis sonic and the static pressure at the plane of the nozzle is greater than the ambient atmospheric pressure. At high speeds this discrepancyis intensified and represents a considerable thrust loss over what could be achieved if the gas was expanded supersonically down to atmos-spheric pressure. For example, at a flight Mach number of 2.0 some- thing like 6 per cent gain in gross thrust is available, and taking accountof the (high) intake momentum drag this is equivalent to a gain of about 18 per cent in net thrust and in specific fuel consumption. Inpractice it may not be possible to achieve the whole of this ideal due to incomplete expansion and irregularities in the nozzle profile but thepromised gains are sufficient to warrant further development. A further advantage is that the divergent part has a considerable effecton reducing the base drag or the afterbody drag of the aircraft. On a typical application the net gain in the value of thrust minus drag atM = 2 was estimated to be 32 per cent. As on most attractive features there are associated disadvantages. In this case it is the attendantloss of thrust when the nozzle operates below the design expansion ratio. This can be minimized by varying the nozzle geometry, either mechanic-ally or aerodynamically, but either method will, of course, demand considerable development. "The indications are that the pure jet engine will still be the power-plant for future high-speed aircraft, certainly up to M = 2.5 and pos- sibly even higher if the developments toward increased flame tempera-ture materialize. With increasing flight speeds the optimum engine pressure ratio will decrease until logically the optimum engine is onein which all of the compression is effected in the intake, i.e., the ram- jet. Its inherent lightness together with ease and cheapness of produc-tion make the ramjet a very attractive powerplant but unfortunately its characteristics are such that some additional form of thrust boostingis required for take-off and acceleration. This drawback is not so serious in the case of guided weapons, since their requirement of highinitial acceleration necessitates some form of thrust boosting in any case. Transatlantic engines "One of the 'plums' of the aero-engine business is that of providingengines for the airlines and in particular, in the near future, we shall have the battle for the first non-stop North Atlantic service. Theproblem of deciding the best type of engine for a duty such as this is very complex. For example, one is apt to regard fuel consumption asof prime importance whereas invariably reduced fuel consumption is only achieved at the expense of increased engine weight. Furthermore,comparisons between different engines are often made by application to the same aircraft, which may be the best configuration for one enginebut will certainly not be the best for all engines. For long-range opera- tion, the decision appears to rest between three types of engine—thehigh-pressure-ratio turbojet, the by-pass engine and the turboprop. "Basically the gas turbine can be regarded as a means of producinggas at high pressure and temperature. The energy in this gas may then be used in a number of ways:— (1) By expanding to atmospheric pressure in a nozzle; this is thesimple turbojet. (2) By partially expanding in a turbine driving a fan, and thenfinally expanding in a nozzle; this is the by-pass engine. (3) By partially expanding in a turbine driving an airscrew andthen finally in a nozzle, giving the turboprop. "It is obvious that, for a given engine operating cycle, and, therefore, for a given thermal efficiency, the jet efflux of the by-pass engine will becooler and at a lower pressure than that of the corresponding pure jet. Thus the by-pass engine will give less thrust but will have a betterpropulsive efficiency and, therefore, a better fuel consumption. A similar improvement in propulsive efficiency of the jet engine can beobtained by operating at a lower flame temperature, but, of course, only at the expense of thermal efficiency. A fairly reliable comparisonof a 'cool' pure jet and a by-pass engine has been made and yields the following results. First, if we choose the operating flame tem-peratures so that each engine is operating near its minimum s.f.c, which is the same in both cases, the installed pure jet will weigh 20per cent more than the by-pass engine and will give 13 per cent more take-off thrust by overspeeding to the same flame temperature, assumingthe same cruise thrust in each case. On this basis the cruising flame temperature of the pure jet would be 120 deg C lower than the by-passengine. "Alternatively, taking the same take-off thrust, flame temperature,and pressure ratio in both cases, the installed weights would be the same, but the by-pass engine would have a 4J per cent advantage infuel consumption at the same cruise thrust, although its cruise flame temperature would be 45 deg C higher. In either of these cases theby-pass engine would show a range advantage of around 5 per cent. This marginal performance advantage is supported by certain installa-tion advantages due largely to the lower skin temperature of the by-pass engine. Turboprop limits "The overall efficiency of the propeller-turbine is largely governedby the efficiency of the propeller, which is dropping rapidly above 450 m.p.h. Fig. 4 shows a comparison between the cool jet, by-passand propeller-turbine engines on the basis of operating costs. This shows the advantage of the by-pass engine over the pure jet at all stagelengths, and further shows the penalty of operating the propeller- turbine at speeds in excess of 450 m.p.h. When one takes into accountthe increased weight, the steadily increasing noise level from the propeller at higher flight speeds and the difficulty of designing forsafety in the event of failure of any part of the complicated pitch control mechanism, it is felt that one cannot make a case for thepropeller turbine above 450 m.p.h. Fig. 4 does, however, emphasize the superiority of the propeller turbine for operation at speeds below450 m.p.h. "Although it is always dangerous to generalize, it seems certain thatthe propeller-turbine will meet the transport requirements up to 450 m.p.h. with the by-pass engine showing a marginal theoreticaladvantage at higher speeds. As has been stated earlier, this marginal advantage can only be consolidated by having a dependable engineof the correct size available at the right time." In the next part of his paper, Mr. Newton reviewed the possibilitiesof "Flying Bedstead" type machines, and discussed the possibilities of scaling engines up and down in size. He finally turned his attentionto fuels, and pointed out that gas turbines had not so far required any parallel line of development to the increase in resistance to detonationneeded by advanced piston engines. For hydrocarbons it appeared that no safe fuel was significantlybetter than any other, on either a weight or volume basis. Boron had about 75 per cent better calorific value than carbon for comparabledensity, and hydro-borons might be expected to be attractive. At a given air specific impulse, hydro-borons were better on both a volumeand a weight basis; the improvement was, however, reduced at high outputs owing to vapourization of the oxide at above 1,800 deg K.The oxide would be liquid at present-day turbine-inlet temperatures, and could therefore deposit or erode. The best application for suchfuels might therefore be in afterburners, below 2,000 deg K. Boron and hydrogen and their compounds were, at stoichiometricconditions, able to give some ten per cent better air specific impulse than kerosine. Values 30 to 35 per cent higher than kerosine werepossible—at the cost of very high s.f.c.—by burning magnesium and aluminium (see Flight for November 21st, 1952). Work was also goingahead in the investigation of metal, and metal/hydrocarbon, slurries.
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