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Aviation History
1954
1954 - 0034.PDF
20 FLIGHT SUPERSONIC SHAPES Lockheed's Chief Engineer Questions the Delta Configuration AT the S.A.E. National Aeronautical Meeting in Los /% Angeles, California, Mr. Clarence L. Johnson (chief engineer of the Lockheed Aircraft Corporation) delivered a noteworthy lecture on the subject of aircraft configurations for flight at between Mach 0 and 2. A summary appears below. Saying in his introduction that he intended to touch on a few factors concerning power plants, but to concentrate chiefly on aircraft configuration, the lecturer remarked that he looked witfi some disfavour on the amount of publicity which had been given to the delta wing form—on which, he claimed, there was little flight experience and in some cases only very limited wind tunnel data. The delta had become fashionable, but, he said, "styles are bad things in engineering." On the other hand, the afterburner, viewed by many as a "contraption of the devil," had come to stay. At a Mach number of 2 die afterburning engine developed 320 per cent of its static thrust, while tfee non-afterburning unit developed only 180 per cent. It was therefore apparent that, for a given thrust at Mach 2.0, an engine with afterburner was lighter, more compact, and easier to install than another,without afterburner, large enough to produce the same thrust at the same speed. A comparison of weight alone showed a 25 per cent advantage to the afterburning engine. Mr. Johnson ruled out rocket power because of its fuel consumption and short endurance, and plain ramjet power because of its limited speed-range. With the help of a diagram (Fig. 1), he showed that engine thrust had increased by 300 per cent since 1945, while airframe drag at Mach 2 had been reduced by 50 per cent. These advances had made possible the design of practical supersonic aircraft. In addition, he stated that high-speed turbojet aircraft attained optimum speed performance at isotropic altitude, i.e., 35,000- 40,000ft. Above this height, the jet engine was no longer favoured by decreasing air temperature, and its efficiency fell off rapidly. Mr. Johnson now turned his attention to aircraft condgurationi, dealing first with the delta wing. Very few delta aircraft had so far reached high speed and, in fact, few attained the speeds of Service aircraft. He saw the advantages of the delta wing primarily in its structural features. The tailless delta was com mitted to flight at low angles of attack because of high drag penalties at high lift settings. Trailing-edge elevators decreased lift considerably. This, together with high leading-edge sweep, produced low maximum lift coefficients—the chief disadvantage of the delta. Even the structural advantages of the delta wing could be found equally in a rectangular wing with an aspect ratio of 2 or 3. Delta aircraft made up for their lack of horizontal tail surface by requiring a very large fin and rudder area. In fact, the speaker demonstrated by a table [opposite] that the tailless delta often had the same total surface area as an aircraft of conventional configura tion. He said that the low wing-loading required for a delta in order to obtain reasonable take-off and landing distances resulted in fact in greater surface area per pound of gross weight than with conventional types. For supersonic aircraft, this was a most undesirable trend. The provision, however, of a normal tail for a delta wing allowed higher wing loadings and improved overall performance. The lecturer claimed diat supersonic flight required wing thicknesses so low that the advantages previously credited to the delta wing disappeared. Since it was impossible to reduce the cross-section of the engine, military equipment, air ducts or pilot THE subject (or title) of this lecture is one of wide interest, and the views of the Lockheed Aircraft Corporation as expressed by its chief engineer are well worthy of study. For this reason we publish the accompanying summary of a recent S.A.E. (Los Angeles) lecture. Some of the facts, arguments, figures and deductions may come strangely to British readers, and those concerned with delta-wing aircraft in this country may here and there disagree—as, no doubt, will Convair engineers. The marked difference between delta-wing considerations for high subsonic and supersonic flight should not be overlooked. by a choice of wing plan-form, the delta wing, witii its ample stowage space, in the end showed no advantage over the normal wing. While fuselage size had to be essentially the same for straight or delta-wing configurations, the delta, with its large fin, merely produced greater aerodynamic drag. While low wing- loading sounded good—"pounds per square foot falling off the slide rule"—maximum lift required very high angles of attack. Mr. Johnson maintained that the delta had limited pitch control and a low usable maximum lift because of limitations in angle of attack during landing and take-off. The speaker next dealt with a 1947 Lockheed design-study for a twin-fin delta fighter. Construction had actually started when the project was abandoned because it "showed so many aerodynamic problems" during extensive wind tunnel testing. Onepf the major difficulties was stability at high angles of attack, the most dangerous being the reversal of directional stability prior to the wing stall. The delta became directionally dynamically unstable immediately above the stall, whereas the straight-wing aircraft became directionally neutrally stable considerably above the stall. Even the use of twin fins did not correct this fault in the delta. The delta's loss of lift at high angles of attack made landing approaches at high power settings necessary, and great difficulty was experienced in trimming the aircraft for such angles of attack, without spoiling its stability. L/D ratio, gliding at a speed 40 per cent above the stall, with 3.7 : 1 for the delta and nearly twice this value for the F-80. The lack of round-out obtainable for landing required long oleo-leg travel which, on the Lockheed design, turned out to be 25in. While the basic wing structure was quite light, it became heavy after stressing for high suction forces on the upper surface encountered during high-g pull-outs at low altitude, and because of the large number of stressed access-panels which had to be included in the wing surface. This delta wing finally became several hundred pounds heavier than an equivalent swept wing for a conventional configuration. The Lockheed design was eventually discarded because of the high drag of the wing over the transonic speed range. The swept wing was found to have good characteristics at high subsonic and transonic speeds. It also had disadvantages, such as high structure weight, aero-elastic problems, low maximum lift, and pitch instability near the stall. For supersonic speeds the swept wing was, Mr. Johnson considered, inferior to straight or delta layouts. A straight wing was defined as one with zero sweep between 25 and 70 per cent chord; it would probably have to be tapered for structural reasons. If a low aspect-ratio was required, it was quite possible to design a straight wing with a thickness/chord ratio equal to that of the delta. The major problems would he in aero-elasticity and flutter, but even these were not insoluble. The straight wing, because it could provide lift coefficients better than those of the delta, even when extremely thin, gave better take-off and landing characteristics than the delta. Conversely, Fig. 1. (Left) Comparison of thrust and drag against Mach number between fighter aircraft 1945 and 1955. Fig. 2 (Right) Effect of thickness on the drag of a straight wing with no leading-edge devices and with a level-flight lift coefficient. SOO 400 °300 u200 < a (OO •6 8 IO (-2 MACH No. I \ ^y T . i 3^T IT / -IICK WING rri T %> -S% *•-* \-A 4-6 <-8 20
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