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Aviation History
1964
1964 - 2052.PDF
FLIGHT International, 9 July 1964 73 recoverable, economically the velocity increment should be as large as possible. For a velocity increment of 6,000m/sec it is three or four times cheaper to have a reusable hydrogen/oxygen stage capable of 20 flights. It is most important with reusable stages that the optimum engine size be used. The move from maximum payload to minimum cost is (for liquid-propellant stages) achieved by decreasing the proportionate engine size and initial stage acceleration. Although the mass at launch increases, the real engine size and cost decreases more rapidly, thus producing the reduced stage cost. When extran- eous costs (some of which will be sensitive to vehicle size) are introduced, the optimum stage will not correspond exactly to the minimum-cost stage. In European space technology it appears certain that boosters will be "workhorses." Through their lives they will be called upon to perform a variety of operations which it is not possible to guess at the outset. To compare first stages at identical velocity increments is therefore the only possible approach. The problem of the optimum use of the stage for various needs can be left to those who have more precise end-points, and a first stage already in existence. In selecting the size or velocity increment to be contributed by a single-shot first stage, there will be a tendency to be somewhat modest. Money in a "small velocity-increment booster" may be considered politically convenient. But, with a reusable first stage, the trend is completely reversed. As the planned number of flights to life increases so does the need to increase the size of the reusable part. In the limit, maximum economy is achieved by not throwing anything away. This can mean one stage to orbit, or separable reusable stages. Hypersonic Ramjets Hypersonic Ramjet Research by R. Hawkins of Bristol Siddeley Engines Ltd, Advanced Propulsion Research Group. The technical feasibility of hypersonic ramjets is well established, at least up to Mach numbers of 7 or 8, but what of the engineering problems? They are severe but by no means PRESENT CAPABILITIES. POSSIBLE FUTURE CAPABILITIES. PRESENT COATING LIMIT/ I ////////// LYBDENUM ALLOYS ^OBIUM ALLOYS/, //////// :KEL ALLOYS./ J. 00 3OO I.OOO I.2OO I.4O0 I.6OO I.8O0 2.0OO 12OO 2.4OO METAL TEMPERATURE — *C Temperature capabilities of some refractory metals (Hawkins) Space launcher envisaged in paper by Hawkins would use turbojets to Mach 2.5 and a ramjet from Mach 2.5 to 7 . A feature of hypersonic ramjets having subsonic combustion is the relative size of capture stream tube, combustor, and exhaust nozzle. The latter is usually big enough to fill the entire base of the aircraft, making possible a complete integration of the engine system within the aircraft—a factor which is vitally important as flight speed increases. The most important phenomenon from which the severest practical problems stem is the increase in ram temperature with flight speed. In stagnation regions on leading edges, and inside the air duct where the air is slowed down prior to fuel injection, surface temperatures could approach 950°C at Mach 5, 1,83O°C at Mach 7 and 2,830°C at Mach 9. Assuming flight times are sufficient to achieve thermal equilibrium, the diagram (page 74) shows that the external compression surfaces of the intake and the majority of the aircraft surfaces, all of which are free to radiate in space, would operate at temperatures well within the capability of present-day steels. The leading edges of the engine, for which blunting must be minimized, would reach tem- peratures greater than those containable with steels unless regener- ative or effusion cooling were used. Inside the engine duct equili- brium temperatures are significantly higher. For high-temperature application, the diagram (left) indicates the most promising structural materials. The equilibrium tem- peratures of all the uncooled internal surfaces of a Mach 7 ramjet would exceed l,800°C so that, on present-day performance, even a refractory metal would be inadequate. For these, surface regenerative or effusion cooling might be employed, using the fuel to extract heat from the wall so that the equilibrium surface temperatures are maintained at acceptable levels. Since the heat sink available in the fuel may be limited in relation to the total area to be cooled, it would be advantageous to coat the surfaces with a ceramic so that the acceptable equilibrium surface insuperable: from the trend of results now being obtained from pilot engineering studies it appears that these problems could be overcome within a reasonable timescale. The paper attempts to outline the most severe problems facing the propulsion engineer and to indicate what research is being brought to bear to find their solution. As an example it is assumed that the ramjet to be examined is the high-speed component of a multi-engined powerplant for a space launcher (illustration above). Turbojet engines would be used for speeds up to, say, Mach 2.5, after which a ramjet would provide the propulsion for acceleration to, say, Mach 7. The ramjet has an intake which is positioned to take advantage of the aircraft compression field. The intake has an internal super- sonic compression and diffuses the airflow to subsonic speeds. It is assumed that kerosine is used as the fuel. Finally a convergent- divergent exhaust nozzle expands the combustion gases. Composite structure in which surface is coated with a ceramic (Hawkins) A f> v. -'/ \', '.' v. /' V, .;•''''' LBASE STRUCTURE TYPICAL SECTION OF COMBUSTION CHAMBER VWLL j USING TROWELLED CERAMIC no i
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