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Aviation History
1950
1950 - 2024.PDF
482 FLIGHT, 30 November 1950 ROCKETS and R.A.T.O. A Digest of Mr. A. V. Cleaver's Lecture to the Roya\ Aeronautical Society ON Thursday, November 23rd, a paper, Rockets andAssisted Take-oft, was presented before the RoyalAeronautical Society, by Mr. A. V. Cleaver, A.R.Ae.S. The lecturer is in charge of the rocket development activitiesof the de Havilland Engine Company. By definition, said Mr. Cleaver, a rocket motor was a type ofjet-propulsion unit provided with means within itself to generate all the material of its propulsive jet. If, as was usual, thisconsisted of the end-products of a combustion process, then not only the source of energy (the fuel), but also the sourceof oxygen with which to burn the fuel and release that energy, was carried along. The oxygen carrier (oxidizer or oxidant)and the fuel were both termed propellants, and might be supplied to the motor in a solid, liquid or gaseous form. Inany case, complete independence of the surrounding atmosphere was achieved. Single compounds, or mixtures in which bothfuel and oxidant were combined, were known as monopropel- lants, and liquid-propellant rockets [with which the lecturewas mainly concerned] could vjse monopropellant systems: nitro-methane, hydrazine, and various organic-nitrates were allexamples of liquid monopropellants. Such systems had not yet found wide application, due to their critical nature, andwhat might by now be termed the classic type of liquid-propellant motor carried its fuel and oxidizer separately, and hence wasknown as a bi-propellant type. The propellants were injected into the chamber, atomized,mixed and burned at high efficiency; thereafter, they expanded through a convergent/divergent nozzle to produce a supersonicpropulsive jet. The minimum cross-section of the nozzle deter- mined the chamber pressure for a given through-put of propel-lants, while the chamber temperature was a function almost entirely of the propellants and mixture ratio used. Lowertemperatures could be obtained by running at propellant-mixture ratios departing considerably from the stoichiometric (" chemi-cally-correct ") value, and such a course was often followed to ease the severe cooling problems encountered. Because ofthese, one of the propellants was usually circulated in a jacket surrounding the chamber before injection on a regenerativecoolant system. If the burning time was short (as in assisted- take-oft units) an uncooled chamber could be used, either oneof high heat capacity, or one constructed of some refractory material. Reaction Processes Chamber pressure exerted only a secondary effect on flametemperature, through the mechanism of dissociation effects —i.e., the tendency of the gaseous products to revert intosimpler forms as their temperature was raised, so absorbing some of their available energy and reducing the final equilibriumtemperature, e.g., H 2O—>-H_, and O2, CO,—^-CO and O2;H 2—>H. These effects were rather less marked at higher cham-ber pressures, so that the intended combustion reactions went further towards completion, and the flame temperature, in con-sequence, was somewhat higher. The lecturer thought it might be as well to qualify the useof the terms "fuel" and "oxidant," and to make it clear that any exothermic chemical reaction which could be made totake place in the chamber could form the basis of a rocket motor. Usually, the reaction employed was the combustionof a fuel in an oxygen-rich atmosphere, but this was not essen- tial. Fuels could be burned with fluorine, for example, whenno oxygen was present—although a rocket engineer would still adopt the convention of referring to whatever fluorine compoundwas used as the " oxidant." Again, extensive use had already been made of low-energy systems, utilizing the catalytic decom-position of hydrogen peroxide (2H 2O2 [liquid)] = 2H2O [gas] + O2+ 380 C.H.U./16)* Here, although oxygen was present, there was no combustion of a fuel with it. The thrust of the rocket motor was extremely high, relativeto its size and weight, for two reasons. The first and more * C.H.U. ^Centigrade Heat Unit. The Boeing XB-47 employs 18,000 Ib of take-off rocket thrust m I addition to the 24,000 Ib of its six General Electric J-47 turbojets. * obvious was the essential simplicity of the mechanism; thesecond was even more fundamental. In a rocket motor, the mass-flow through the system was small for a given thrust,and, moreover, a high density of working fluid was maintained throughout. All pumping operations were carried out with thefluid in its liquid phase (the analagous argument for the solid rocket was self-evident), and liquid pumps were inevitably muchsmaller than, for example, air compressors. Once the gaseous phases of the reaction were attained, the rocket motor wasnormally working at chamber pressures of at least 300 Ib/sq in (at all altitudes) as compared with the 60 Ib/sq in typical ofthe gas turbine at sea level (and this was a figure which decreased with altitude). The high specific consumption of the rocket motor was, of .course, unfavourable, and even the most enthusiastic of rocket advocates would not suggest diat rocket propulsion could beemployed other than for short durations, simply because of this factor. Since the heats of reaction of all possible chemicalcombinations were by now well known, it was idle to hope for ' any sensational discovery of a natural " super fuel." Indeed,the hydrogen/oxygen combination was usually regarded by rocket engineers as a sort of chemical ideal. Any sweeping advances must come from the field of inven-tion, or at least from nuclear rather than chemical discovery. It was conceivable that future research into the stability ofmatter might make it possible to manufacture synthetic highly- exothermic propellants, or to use monatomic hydrogen (re-combining into its molecular form) as a monopropellant. More probably, it might be possible eventually to use nuclear reactorsto supply heat to the working fluid (hydrogen, ammonia, or some other light gas) of rocket motors. All these were long-term possibilities, and were certainly of no application to the immediate aircraft case. Nor were they necessary to it; the •correct attitude at present was surely to make the best use of ' -. existing propellants, in order to realize the potentialities peculiarto rocket propulsion. Progress in detailed engineering would v enable the present performance optima to be improved—perhapsby as much as 10 per cent—but, broadly speaking, the limita- tions imposed by high specific consumptions had to be accepted. In dealing with the choice of propellants, Mr. Cleaverremarked on the high cost of solid propellants, and cited cordite, which had been in large-scale production for many years, and rhad been the standard British solid propellant so far: it cost : '
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