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Aviation History
1955
1955 - 0022.PDF
22 FLIGHT,\1 January 1955 WING DROP and PITCH-UP Some Phenomena of High-speed Flight Discussed in S.A.E. Lecture ONE of the papers read at the Los Angeles meeting of theAmerican Society of Automotive Engineers was byEarle S. Hodder, aerodynamics supervisor, North American Aviation, Inc., who dealt with some of the problems of high-speed flight. Mr. Hodder made it clear that his paper was a review of some of the more important problems which had arisen in this field since World War 2, and which could be dealt with without infringing security regulations. His opinions were not necessarily those of his company. Flow Characteristics. The speaker devoted the first part of hislecture to a review of the changes in airflow about "an aerofoil as a function of Mach number. He showed a number of diagramsillustrating typical flow configurations over an aerofoil; these, he said, were substantially the same for any aerofoil, althoughthose with sharp leading edges and small thickness were less seriously affected. Compression shock-waves occurred when the flow over a por-tion of the aerofoil reached sonic speed, and formed at the point where the flow once more became subsonic. The accelera-tion occurred smoothly. Waves normally formed on the upper surface before they did so on the lower surface, and as the speedof the flow increased above Mach 1.0, the wave was progressively inclined rearwards. A bow wave ahead of a wing formed in asimilar way. Another important characteristic was the separation of smooth flow from the wing skin close to the base of a shock-wave. Over a thick wing, the separation was considerable, but it was much less on an extremely thin wing. This area of separationfluctuated over the wing surface and caused a stalled condition which was responsible for such effects as buffet, and controlsurface buzz or tap. Mr. Hodder quoted some measurements taken of the widewing-wake resulting from shock-waves and flow separation at the transonic drag-rise for the F-86D. At a Ci,of 0.25 (equivalent tolg at 40,000ft) an increase of Mach number from 0.9 to 0.95 in- creased the wake thickness by only 30 per cent, while at a CLof 0.55 (2.3g) the wake thickness increased about 100 per cent with the same Mach number increase. Further disturbanceswhich gave some measure of the energy lost in the wake might, for the two above cases, show even greater increases with risingspeed. Wing Drop. Having thus stated the basic situation, Mr. Hodderwent on to deal with wing drop which, occurring near the drag rise, was a result of the compressibility effects just des-cribed. Normally shocks occurred on each wing as an aircraft entered the transonic range and were accompanied by flow separa-tion. Inequalities in the wings, originating in manufacture or subsequent wear, would cause the resultant loss of lift to occuron one wing before the other, with a consequent wing drop or "roll off." Opposite aileron tended to aggravate this, since thedown aileron increased shock intensity with further loss in lift. An aileron might in any case be operating in the separated-flowregion and therefore have a greatly reduced effectiveness. Aileron deflection further aggravated wing drop by twisting thewing, particularly at low altitude, or on a wing of a low stiffness. The speaker quoted the case of the F-86D, where the aero-elasticeffects of aileron deflection made it impossible to raise the wing •90 98 102 TRUE MACH NUMBER by aileron at high speed below 5,000ft. Individual aircraftvaried, of course, some dropping the left and some the right wing, and some having no wing drop at all. Decreasing the F-86D'sT/C ratio from an effective 9.2 per cent to 6 or 7, the speaker thought, would undoubtedly minimize wing drop and even elimin-ate it. To achieve similar results with a straight wing would require a maximum T/C ratio as low as 5 per cent. Such low ratios might free the supersonic fighter from wingdrop troubles; but in bombers and transports, though the ratios might be the same, the trouble might still occur if the wing struc-ture were designed to lower stiffness and strength factors. This would particularly aggravate aileron deflection characteristics. Yet another factor affecting wing drop in the supersonic aircraftwas thought to originate from the bow wave from the nose of the aircraft, a wave which was substantially the same as its counter-part on the wing. As Mach number increased, the bow wave in- clined ever more sharply rearwards, and any asymmetry in itcaused by slight yaw or structural inexactitudes might cause it to affect one wing before the other. Asymmetric shock waves from MINIMUM STALLING SPEED 100 800 900 Fig. 1. Aileron deflection required to hold the wings level in the F-86D at various altitudes. The sea-level line can be seen to exceed the "maximum available deflection" line. 200 300 100 500 000 700 MAXIMUM LEVEL FLIGHT SPEED (kt) Fig. 2. Comparison of maximum level flight speeds and stalling speeds tor a wide variety of American aircraft since 1906. The stalling speeds are those for the landing configuration. other parts of the structure forward of the wing might also giverise to wing drop. Pitch-up, Dig-in or Overshoot. This effect, the speaker said,was first encountered in World War 2 fighters and had occurred again from time to time. It consisted of a sudden tightening of aturn or pull-up to a g-loading considerably higher than the pilot had intended to apply. There were several causes, and one ofthe earliest had been the elevator trim-tab. The pilot would begin a dive and would progressively trim-out stick forces with theelevator tab. Eventually, however, shock-waves and flow separa- tion on the tailplane would drastically reduce, if not cancel out,the effectiveness of the tab. Because of high stick-forces, the pilot would then use the trimmer during the pull-out, although it wasineffective; but, as the load factor increased, the aircraft slowed down until, as the sho;k-waves dissipated, the trim tab wouldregain its full effectiveness and cause a rapid pitch-up which the pilot would not be able to control. The speaker stated, however,that this was not likely to occur nowadays, since no one in his right mind would now use a conventional elevator-tab system ona supersonic aircraft. Another cause of pitch-up was the large increase in staticstability margin with rising speed. If a pilot flying at Mach 1.2 and a CL=0.1, pulled up to 5g (CL~0.5), the increased dragwould decelerate the aircraft to perhaps Mach 1.0. While with increasing speed, the c.p. would move back over the wing,as the speed decreased again it would once more move forward; and if the tailplane angle were maintained as for the originalg application, its effectiveness would increase with decreasing speed. The result would be a nose-up pitch. This characteristicwas already experienced in transonic aircraft and would doubdess continue to occur in supersonic aircraft. It was aggravated by therelatively low effectiveness of the elevator at transonic speed. The effects could, however, be overcome to some extent byvery powerful tailplane controls; in any case, the movements of centre of pressure would not be great above Mach 1.5. Another source of pitch-up was found between Mach 0.6 and0.9 at a C L between 0.4 and 0.6 Pitch-up might then occur
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