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Aviation History
1950
1950 - 2025.PDF
FLIGHT, 30 November 1950 483 about £280 per ton. As shown in the curves, the empty weightof the installation mattered most for assisted-take-off units. The weights before firing could be derived by adding on tothese empty weights the propellant charge weight given by: thrust/specific impulse x burning time. For the liquidpropellant motors, the curves illustrated the superiority of the pump-feeding system for longer burning times. As regards liquid propellants, the first essential was to realizethe very wide choice available—in theory, from among any exothermic chemical reactions. So far as bi-propellant systemswere concerned, it should be realized that, in general, any fuel could be burned with any oxidizer, although, in practice, theselection of a novel combination could involve considerable development work to solve problems of ignition, smooth andefficient combustion, and so on. On aircraft installations, there was an obvious argument for using conventional fuels (petrolor kerosene) and there was little to be gained performance-wise by employing more unorthodox fluids. The case for unfamiliar fuels arose mainly in connectionwith the problem of ignition. Certain risks incurred during starting were thereby avoided, but only at the expense of intro-ducing others in connection with storage, handling, crash con- sequences, and so forth. It was the lecturer's opinion that, onbalance, such combinations were not the best choice for aircraft use, although they might well be so for some missile applica-tions. The main choice rested between the many possible alternatives for the oxidant, but in practice, at least for thepresent, these reduced to the three fluids, liquid oxygen, nitric acid (99 per cent strength), and hydrogen peroxide (80 per centstrength). Each of these had its merits and disadvantages, and the same lack of any universal superiority of one or the otheremphasized the fact that the choice could be made only for a particular type of application. Considerations of personal judg-ment and experience would also influence the choice as they always did in matters of engineering compromise. Mr. Cleaverfurther gave it as his purely personal assessment that the fugitive nature of liquid oxygen made it unsuitable for aircraft applica-tions, and for such duties, either hydrogen peroxide or nitric acid (in that order) were preferable. Detail Design The general engineering problems were concerned chieflywith the demands made upon the large number of hydraulic and/or pneumatic valves, automatic and otherwise, which wereinvariably necessary in rocket motors. The design of these components was complicated by the fact that (a) large fluid-flowrates had to be handled; (b) the fluids concerned might have peculiarly different properties, e.g., corrosion (nitric acid), lowtemperature (liquid oxygen), or liability to decomposition (hydro- gen peroxide); and (c) unpremeditated mixing due to leakageof highly energetic fuel combinations had to be prevented. This was particularly necessary with spontaneously-igniting propel-lants, but also required attention in other cases. Although an intimate mixture or emulsion of fuel and oxidant, in a dangerousratio, was unlikely to be produced accidentally, the risk was potentially present and had to be avoided. Heat transfer represented a physical effect imposing morearduous design conditions on the rocket motor than in any other engineering application. The Walter motor for the Mel63,for example, ran at a relatively low chamber temperature (about 2,000 deg K) judged by rocket standards. Nevertheless, theheat transfer at the most critical part of the nozzle was still at the rate of over 200 C.H.U./sec/sq ft. This most critical position was always near the qpzzle throat,where the stagnation temperature at the boundary layer was still near to the bulk gas temperature upstream in the chamber,while the high gas-flow velocity eroded the thickness of the boundary layer until it provided only a thin insulation for thechamber wall. The use of internal " film " or " sweat" cooling, in addition to the normal regenerative system, operated toalleviate this condition in two ways. First, the evaporation of the film coolant itself provided additional local cooling; secondly,it restored a thick boundary-layer insulation. Such a technique could be applied (as on the V-2) by introducing one of the a O12 x t~ * t 008 2 ui O O6 - OO4 * 002 § ° UJa. Ml 5S-- \ hJ/^—LIQt r I OLID S / h-H JIO (p RES LIQUID (PUMP-FED<h-i SURE -FEt . 20 4O »O 8O BURNING TIME (tec) IOO 12C Curves of specific weight of aircraft rocket power-plants as a function of total operating period. The values are applicable to motors of circa 5,000 Ib thrust, and include pump and tank weights, etc., but not mounting parts or cowlings. propellants through discrete holes, or by making the chamberwall itself of some sintered porous material. The latter method had not yet found much application because of constructiondifficulties. The choice of chamber materials was governed by the desirefor two conflicting physical properties, good strength at elevated temperatures, and a high thermal conductivity. Thus, therewas the paradoxical situation that ordinary mild steels, high- grade heat-resistant alloys, and aluminium alloys had all beensuccessfully employed on rocket motors operated at tempera- tures of over 2,000 deg K. For uncooled, short-period motors,the use of graphite liners, ceramics, or the new metal-ceramic compacts (" ceramals" or " ceramets") was of interest, as analternative to using, for example, thick copper walls of high heat-capacity; the main difficulties to be overcome were those ofconstruction of components in useful shapes and sizes, and of improving resistance to thermal shock when the motor wasstarted. Mr. Cleaver then gave a brief survey of the history of modernrocket work, in which he stated that Dr. Johannes Winkler was financed by the Junkers concern in 1929 to conduct experimentsin rocket take-off assistance for seaplanes. British interest in rockets for take-off assistance first became pointed with thedevelopment of small, solid-propellant (cordite) rockets in 1940, and very little British work was done on liquid-propellantrockets until after the war. During the past two years, however, de Havillands had developed a large "cold" peroxide assistedtake-off unit called the Sprite*. The "cold" peroxide system had been selected because of its simplicity and the large Germanbackground of successful experience with it; moreover, since the possibility of using the units had been foreseen at an earlystage, they had been designed for internal installation in the wings of the Comet. The Germans, however, were developingliquid types even before the war, and one of the early projects of the great Peenemiinde establishment was a petrol/liquidoxygen unit for assisted take-off. Fairly extensive research on both solid and liquid rockets wasin progress in America even before 1939. The Reaction Motors Corporation, formed by a number of the leading members ofthe American Rocket Society, developed several assisted-take- off units, and latterly had produced the power-plants for theBell X-l and Douglas Skyrocket research aircraft and for the Convair 774, the Martin Viking and other missiles. The Gug-genheim Aeronautical Laboratory, California Institute of Tech- nology, was also very prominent in this field, and the AerojetCorporation was formed to apply the results of their research. Among the many rocket motors produced by this firm therehad been a number, both of solid- and liquid-propellant types, designed for take-off assistance. The liquid-propellant designs * See "Flight," February 2nd, 1950. The de Havilland Sprite — o large "cold " peroxide assisted- take-off unit. Provision for two Sprites is made in the Comet. The "cold" peroxide system is simple and is backed by a wealth of German experience.
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