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Aviation History
1946
1946 - 2120.PDF
444 FLIGHT OCTOBER 24TH, 1946 HIGH-SPEED AIRFLOW Fig. 12. TheLippischP.n. use of tailless or tail-first layouts and triangular wings orwings of low aspect ratio, were suggested at that time. The German scientists were aware of the increase in longi-tudinal stability, but their attempts to minimize it were very empirical in character—a most unusual event in German aero- nautical research, probably a consequence ofthe urgency with which the solution was wanted for guided missile application. Thusat Kochel, the major problem in the develop- ment of the supersonic guided missile" Wasserfall " was to achieve a roughly constant " neutral point " at subsonic andsupersonic speeds; and this was done by test- ing a whole series of wings and tailplaues,some of which are shown in Fig. 15. The final version, third from the left in Fig. 15,had a neutral point movement of less than 0.03 chord over the whole range of Mach num-ber from o to 2.5. The change in wing plan- form from the original version at the right-hand side is noteworthy. A similar series of experiments on therocket, A.9, which was a winged version of the V.2 rocket designed for much greaterrange, showed the same difficulty in obtain- ing satisfactory longitudinal stability aboveand below the speed of sound, and in this case no really satisfactory solution was obtained. Before leaving the subject of wings of small aspect ratio,the large amount of work on their low-speed characteristics TRIANGULAR AEROFOIL CHARACTERISTICS AT LOW MACH NUMBER Fig- Aspect raik> Apex angle (degrees)y Aerodynamic centre h A.C. at supersonic speed* i 3.82.7 X125 0.33 253.2 2.1 0.185 0.33 4/336.8 1.6 0.19 0.33 128 1.23 0.215 0.33 * Kougbly at supersonic speed the aerodynamic centre is at the centre of area ; this, compared with the value of h at subsonic speeds, indicates the longitudinal stability change. should be noted. A series was tested at the D.V.L., and slopeSC oi the lift curve —- and aerodynamic centre distance h be- hind the mean quarter-chord pointwere measured. A few results on tri- angular wings are given in the table.It is of interest to note that results confirmed Lippisch's statement thatno wing-dropping tendencies appeared at aspect ratios below about 2J; andthat maximum lift coefficients of about 1 o were obtained with aspectratios of 1 to 2, but at incidences of nearly 40 deg. The superiority oflow aspect ratio triangular wings in reducing the stability change attransonic speeds is apparent from the last two lines of the table. Much of the supersonic work inGermany during the war was, of course, aimed at developing pro-jectiles and guided missiles, although in many instances the results pointthe way to future supersonic aircraft, as indicated in the paragraphs above.An Institute in Vienna under Dr. Lippisch was studying the supersonic 0-05 0-04 0-03 0-02 0-01 Fig. 14. Drag of Lippisch P 13 aerofoil at high speed. aircraft, however, and the proposals shown in Figs. 12 and 13were originated here. In addition, the D.F.S. (Deutsch Institut fur Segelflug) at Ainring was constructing an experi-mental aircraft powered by two bifuel rocket units, intended primarily for aerodynamic work at Mach numbers up to 1.4(although estimates of its performance indicated a maximum Mach number of 2.7). A wing of 35 deg. sweep-back and normalaspect ratio was proposed, with a tail of conventional type; apparently the intention was merely to accept the longitudinalstability increase. In fact, it appeared that much more ingenuity had been expended on problems of pilot's escapethan on aetodynamics; possibly a wise step! Shock Diffusers The problem of efficient conversion of kinetic energy topressure at supersonic speeds is, of course, of major importance to turbo-jet or ram-jet power plants for supersonic speeds,when the pressure ratio obtainable by slowing the stream is sufficient to give quite an efficient thermodynamic cycle. Theproblem also has an important side-line in the design of super- sonic wind tunnels. A major contribution in this field was made by t>swatitsch, r.of Gottingen. He suggested that the normal shock which appears in the entry of a duct at supersonic speeds could bereplaced by several oblique shocks, across which the compression of the air would occurmuch more efficiently. The practical method of achieving this, by means of a central corein the entry, is shown in Fig. 16. The position of the outer lip relative to the noseis chosen so that the nose wave impinges on it, and further oblique shock-waves arelocated by corners in the profile of the central core. Several oblique shock-waves in seriesalong the entry duct, ending with a final normal shock of small strength, are indicatedin Fig. 16; the greater the number of oblique shocks, the better is the pressure recoveryin the diffuser. This is demonstrated in Fig. 17, which is extracted from recentBritish work, and compares the pressure recovery pS (pS is the ratio of the actualpressure increase in the diffuser to that which would be obtained in isentropic flow, i.e.,without shocks) for the Oswatitsch type of difluser and the usual type of pitot entrywith normal shock-wave. It will be seen that the gain in pressure, at Mach numbers above2, is very large with two or three oblique shocks; above this number the further gainsare not great. These theoretical expectations were confirmed in Germanyby supersonic tunnel tests at L.F.A. and A.V.A. Fig. 18 shows the results of A.V.A. tests at ilf = 2.9. The maximumpressure obtained was just over 0.7 of the isentropic ideal, whereas at this Mach number a pitot tube records a stagnationpressure of only 0.35 of the isentropic value. The significance of this change will be appreciated when it is realized that apower unit working in the duct would be accepting air at 20 atmospheres pressure in the first case, and only 10 atmo-spheres in the second. The only practical difficulty with diffusers of this type is toe-sign them satisfactorily for a wide range of Mach number. The angle of the shock-waves, of course, changes with Machnumber, and at Mach numbers well away from the design value the shock-waves from the corner no longer meet the outer lip. There is then a lossof compression efficiency. The ob- vious remedy of altering the diflusershape with the Mach number present?s^S| quite a complicated mechanical *problem Every worker in the high-speedfield very soon comes up against the major problem of theoretical or ex-perimental work in the immediate vicinity of the speed of sound. Welliiefore the speed of sound is attained in subsonic wind tunnels, the flowceases to represent that of the model in free air and the whole tunnel be-comes in effect a convergent-divergent nozzle. This limits the useful Machnumber to a maximum of 0.85 to 0.9. Approaching from the supersonicside, a parallel difficulty appears; the reflection of the nose-wave of the The Lippisch model. P13
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