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Aviation History
1963
1963 - 2242.PDF
0 18 - 014 010 ^Sgg. AIR ^^ INTAKE- PRESSURE LIMITATION 1 i£t*Q „ > ^^~HEAT EXCHANGER METAL TEMPERATURE LIMITATION FLIGHT International, 26 December 1963 1041 6 7 FLIGHT MACH NUMBER Fig J Theoretical minimum fuel/air ratio attempt has nevertheless been made, and yields a value of 3,0001b for the matrix only on lOOlb/sec air consumption. With a = 0.15 and a fuel specific impulse of 900sec, this would correspond to a net thrust of 13,5001b, i.e. to an aircraft weight of 80,0001b at L/D = 6. Adding something for the air intake, pipework and headers, the weight penalty might rise to perhaps 4,0001b, i.e. 5 per cent of the aircraft weight. With a total (fuel + engine) weight of, say, 40 per cent, it is clear that the liquid air cycle engine can well afford to give away a 5 per cent weight penalty if the fuel specific thrust is going to be doubled. When compared with the ramjet, however, the liquid air cycle engine looks much less attractive. The ramjet, with its ability to burn its fuel at stoichiometric mixture ratio instead of extremely rich, can produce a fuel specific impulse of 2,000 to 4,000sec over the speed range from M = 3 to 8 with an engine specific weight varying between perhaps 0.0025 and 0.3, depending on efficiency and flight condition. These figures need comparison with about 900sec impulse and estimated specific weight of about 0.32 for the liquid air cycle engine, or 450 and 0.02 for the rocket. It appears that LACE systems will only compete with the ramjet as a main power plant in the speed range M = 4 to 8 if the problems of the ramjet prove rather less tractable than is expected, so that it becomes necessary to accept various limitations and penalties which could eat into the marked performance and weight superiority of the ramjet. There do exist, however, some possibilities whereby a LACE system could ease the mechanical problems of the hypersonic ramjet and increase its flexibility, the first by reducing the variability required of the engine components, the second by the ability to store power. It may, for instance, be combined with a hydrogen- burning ramjet, taking its air from a common intake and diffuser. Assuming that the liquid air cycle can now make use of the ramjet's liquid hydrogen as well as its own, the disadvantage of having to work on a grossly over-rich mixture is now removed. This arrange ment also has one or two advantages on the ramjet side. The liquid air cycle retains to some extent that important characteristic of extra thrust availability for manoeuvre which is inherent in rocket propulsion. This may be achieved by pumping extra liquid air into the combustion chamber from an accumulator. It is also possible to think of the liquid air cycle as providing a convenient means of lateral flight control at high speeds. The Oxygen-condensation System in Space Vehicle Launchers Prediction of the performance of air-breathing propulsion and lifting systems as applied to space- vehicle launching is severely restricted by the scarcity of real engineering data, on both engines and aircraft structures, for vehicles capable of operating within the atmosphere at flight Mach numbers between say 5 and 15. Lane1 in his paper to the RAeS used assumptions which allowed him to calculate that a hydrogen fuelled air-breathing aircraft launching a two-stage hydrogen/oxygen rocket at a flight Mach number around 12 could Put into a 300 mile circular orbit about 13 per cent of the all-up height. Work at NGTE2 with more severe assumptions suggests a separation Mach number around 7 and a useful load of about 55 per cent. However, considering these latter figures, partly because the demands for advanced technology are less extreme than with Lane's results, we find that some 32 per cent of the weight °f the whole assembly at take-off is due to the liquid oxygen in the tanks of the upper-stage rockets. A tremendous amount of work is done inefficiently during the early acceleration. If this oxygen were not present at take-off, then not only could the work be done at a more efficient stage of the flight—at, say, M = 6—but the induced drag during the acceleration could be reduced and the tanks in the rockets could be used to carry some of the hydrogen used in the air-breather. Against this advantage must be set the weight of the liquefier and separator equipment together with the supplementary air intake and the necessary control gear. The problem of separating the oxygen and nitrogen is discussed later. For the purpose of preliminary project sums, it has been assumed that the weight of the separator gear and any necessary power supplies are equal to the heat-exchanger weight and that a hydrogen/oxygen ratio of 0.2 has been needed in the heat exchanger. The validity of the first assumption is, of course, open to consider able doubt and is only suitable for "feeler" calculations. The second is a little severe for likely operating conditions around 500kt and M = 6, but includes an allowance for the fact that the nitrogen will not, in practice, be rejected at ram air temperature. Since the weight of the liquefier and separator gear, etc, is depen dent on the rate of oxygen storage rather than the total amount collected, it is desirable to collect the oxygen as slowly as possible. This implies a cruise phase in the launching vehicle which, of course, consumes fuel. With the above weight assumptions, which can be expressed as a LACE weight of 3001b for each lb/sec of oxygen storage rate, the optimum occurs with a cruise duration of about 8min, a flight Mach number at which to collect of about 6 and therefore a cruise range of some 500 miles. The weight of the gear is then about 12 per cent of the all-up weight, while the extra fuel carried for the cruise phase is about 3.5 per cent of the all-up weight. Most of this fuel can be carried in the tanks of the upper rockets which would otherwise contain oxygen; such dual-purpose tanks would require a suitable isolation or purging technique, which does not seem outrageously impractical. The percentage weight put into orbit can be increased by this technique to about 7.6 per cent of the all-up weight, about twice the corresponding figure for an all-rocket system. The total size of the air-breathing vehicle would not be much altered, but its all- up weight would be some 17 per cent lower, with consequent advantages in airfield and low-speed performance. However, the more important aspect of such operations is likely to be not the all-up weight of the vehicle, but its research, development and production costs. Obviously these cannot be accurately predicted, but a first-order guess suggests that, since equipment is more expensive than fuel, the total cost of the vehicle including the oxygen-condensation system would not be much different from the simpler but heavier vehicle without it. The oxygen-condensation seems worth some further investi gation with a view to possible application to space vehicle launchers, the most important problem being the achievement of an efficient light-weight device for separating the liquid oxygen from the nitrogen. It must, however, be emphasized that no practical solution is yet in sight in terms of the current state of technology. 1 R. J. Lane, Recoverable Airbreathing Boosters for Space] Vehicles, JRAeS, June, 1962. • R. A. Jeffs and Mrs G. C. A. Jones, unpublished work at NGTE. Fig 4 Theoretical thrust for LACE z O i- CL r> if) 2 O IOOO 800 i- ^ z\ 3 <D a. I 600 Z 400-* I » 6 7 FLIGHT MACH NUMBER.
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