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Aviation History
1953
1953 - 1525.PDF
20 November 1953 679 Even more serious was the limitation of ground angle. On low-aspect- ratio swept wings, large incidences had to be used for take-off and landing, and a long fuselage would then necessitate a very long under carriage with the obvious associated structural and stowage problems. Increasing fin size would improve the oscillatory stability and have only a small effect on spiral stability. The dihedral would usually be just adequate to give spiral stability on the cruise, which might mean spiral instability either at high CL values or at Cx =zero. The provision of sufficient inherent stability had proved difficult, and many aircraft were using artificial stabilization of the oscillatory mode, by the installation of a yaw damper. Comparing the tail volumes of the Comet and the Constellation, their gross horizontal tail volumes were 0.486 and 1.12 respectively, and the net vertical tail volumes 0.023 and 0.064. If the Comet had nacelles and airscrews, then both vertical and horizontal tails would have had to be considerably larger. The combined increase in area would be about 65 per cent; if the Comet had needed as large a tail as this, the associated weight and drag increases would have reduced the payload for a given range by about 18 per cent of the capacity payload. This illustrated forcibly the large gains in performance to be obtained on a jet-propelled aircraft by keeping the tail as small as possible. Turning to tailless aircraft, the lecturer stated that the damping-in- pitch of this type would apparently be increased—certainly at subsonic speeds—if there was a large wing-chord. This was achieved on the low-aspect-ratio delta configuration, which might then provide a good deal of the basic damping which the earlier tailless aircraft of higher aspect-ratio lacked. Low e.g. travel as a percentage of the mean chord might not be so serious on a delta aircraft, as it represented a sufficiently large actual dimension. Three radically different wing plan-forms were being evolved to meet the need for long-range high-speed aircraft. These were: (i) the straight or slightly cranked swept wing of high aspect-ratio, sub divided into those with buried engines and those with pods; (ii) the crescent wing; and (iii) the delta wing. Performance considerations played a large part in deciding on these plan-forms and were mainly a question of compromise between high lift, low high-speed drag and structure weight. Stability and control considerations also had an important influence. Low-speed Behaviour At low speeds, the stall was perhaps the major hazard in fixed-wing aircraft. It could not be eliminated, but it could be tamed. To do so, it was necessary to avoid large rolling and pitching displacements, provide adequate control and ensure warning of the approach of the stall. In the past, lateral instability with wing dropping had always been the main problem. Swept-back wings tended to have a tip stall which, if the wing was also of high aspect-ratio, could cause a nose-up pitching and pronounced longitudinal instability; this could then be more serious than the wing-dropping tendency. In the thin, swept-back wing was one of the most severe clashes of requirements, as the better their high-speed performance the greater the difficulty at low speeds. There were a large number of parameters which affected the stall: wing section, wing sweep, aspect ratio and taper ratio, Reynolds number, Mach number, roughness and flaps; and there were an equally large number of devices available to deal with it—slats, nose flaps, drooped noses, fences, boundary-layer control, and so on. Thus, although the problem was complicated, there was a variety of solutions, as was obvious in comparing the crescent wing with the delta wing. The crescent wing aimed at preventing a tip stall in the usual way, but the delta wing appeared at first sight to contradict all normal principles of stall control, since the stall began right at the tip. This, presumably, was acceptable because only a small area of tip was involved, and the spread of the stall inboard was gradual. Also, the delta wing had a low aspect-ratio and it was well established that the tip stalling caused hardly any longitudinal instability on a swept-back wing of sufficiently low aspect-ratio. The Comet used two forms of spanwise stall control, a built-in increase of nose-radius towards the up (the primary control) and the secondary control, developed by model flight test, of the leading-edge fence. It had been found that the chordwise extent of the fence might be adjusted to suit the section's stalling characteristics, i.e., whether a leading-edge or trailing-edge stall. In the lecturer's experience, fences had been a most effective and flexible device for adjusting stalling behaviour in the prototype flight testing stage. On the D.H.I08, tip slats were effective but full-span slats brought fairly severe wing-drop —which, however, could then be cured by fitting two fences. On the Boeing B-47, the pod engines were reported to prevent the tip stall, presumably because they reduced the local incidence of the airflow above them. If tip stall was prevented, then the ailerons would still retain some effectiveness at the stall. At the same time, however, there were other factors which tended to make the ailerons less effective at low speeds. At high speeds, the rate of roll might be calculated with sufficient accuracy by assuming pure rolling motion only, while at low speeds the associated yawing response was very important. It arose from three causes: (i) the adverse yawing moment produced by aileron deflection; (ii) that caused by the response in roll; and (iii) the adverse side-slip induced by the angle of bank. As the response to aileron on the approach often determined the aileron size, it would be very desirable to eliminate the adverse yaw. This could be done by linking the rudder in with the ailerons, so that sufficient rudder was applied to counter the adverse yawing moment. This should enable a reduction in aileron size or improvement in aileron control to be obtained. Spoilers had also been suggested to help to solve this problem. As at low sneeds, the swept-back wing was liable to tip-stalling troubles at high Mach numbers. The methods of curing it, however, would differ, the emphasis shifting from the leading edge to mid-chord p.wlnS to the presence of shock waves. Thus, improved flow was most likely to be effected by making the wing thinner, but other devices designed to re-energize or direct the boundary layer were possible— NOS E U P 1 IM 1 NOS E DO W DESIRED MAXIMUM CRUISING SPEED V< ^ : P" 1 UNSTABLE TRIM CHANGE H 07 0-8 0-9 MACH NUMBER-M Fig 2. Change of trim with Mach number. namely, vortex genera tors ahead of the maximum thickness to re-energize the boun dary layer, and fences to direct it. Because of the funda mental changes of flow at supersonic speeds, it would probably be necessary to accept an eventual nose - down trim-change in the transition from high subsonic to supersonic speeds. It would be a great asset, however, if this trim-change could at least be postponed beyond the cruising speed range. As the maximum cruising speed should not be much above the drag divergence Mach number, this should not prove an impossible aim. Analysis of existing data showed that some aircraft did in fact have a small nose-up trim-change at about the critical Mach number, which was an ideal characteristic, giving an automatic recovery from any inadvertent speed-rise. On the other hand, some aircraft went pro gressively nose-down and started doing so at Mach numbers which might be very near, or at, the maximum cruising speed. It appeared that txiis latter was the more fundamental behaviour and more likely to occur on a clean aircraft, whereas aircraft with nacelles or large bodies in extreme high- or low-wing positions might have a temporary kink in the CM—M curve (dotted line, Fig. 2). One of the biggest control problems today concerned the form of control to be used at transonic and supersonic speeds. Theory indicated that at supersonic speeds the flap-type control became relatively ineffective, as it could produce lift only on its own surface and could not induce life on the main surface ahead, as at subsonic speeds. Aero-elastic considerations were also very important. Inadequate wing-stiffness might lead to wing flutter, loss of aileron control (aileron reversal), wing divergence, and loss of longitudinal stability. To avoid these troubles, the prime requirement was for a certain torsional stiff ness, but on swept wings flexural stiffness was also important. On a straight wing, aileron reversal speed depended directly on the torsional stiffness, while on the swept wing the flexural flexibility had an additional adverse effect. The requirements were for this speed to be at least 15 per cent above the dive speed, and for a straight wing this usually meant rather more stiffness than for flutter prevention. On the swept-back wing, die necessary flutter stiffness tended to decrease, but the aileron-reversal requirement became more difficult to meet. Spoilers appeared to be one obvious answer to this problem, as they greatly reduced the twisting moment applied to the wing. Another solution often suggested was to use an all-moving tip control. A com promise solution of having a diagonal hinge-line, however, appeared attractive; a small area of all-moving rip tapering off rapidly to a small- chord aileron inboard. Concerning the problem of power boost versus spring tabs, it appeared that, once the mechanical problems of a power control were faced, they were preferable on balance to spring or servo tabs. The main advantages were the small control sizes, the greatly reduced flutter problem, and the relative ease of development in the prototype and early production period (well shown in the case of the Comet). In conclusion, the lecturer suggested that the following were the outstanding problems of stability and control faced by the designer:— It was still not possible to formulate all the standards of stability and control which corresponded to the pilot's judgment of good handling qualities. This problem extended into the design of artificial feel systems, the make-up of good dynamic stability, and of good dynamic response to control application. The need for more knowledge of the aerodynamic forces had extended into the subsonic range, and more information on oscillatory derivatives was urgently required. Flow separation was at the root of many troubles, particularly the low-speed stall, the high-speed stall and buffeting. Also, it influenced many others, such as transonic trim-changes, loss of damping and control effectiveness, and the repeatability problem. One looked hope fully at boundary-layer control as the future solution. Automatic stabilization was here to stay. If designed into the aircraft from the sjart it could lead to a better overall performance. It was not easy to decide where 'to put the tail on a high-speed air craft, and it would be valuable to know whether it was essential for transonic damping on a supersonic aircraft. The main problems of power boost centred on the detailed mechanical design of the control circuit and booster valve, and on deciding what form the artificial feel should take. As a counterpart to the all-moving tail, it would be interesting to see an experiment on an all-moving wing tip in partial or complete form. Some full-scale tests on the spoilers on a high-speed aircraft would also be very valuable. In the discussion which followed the lecture, several questions referred to the "feel diagram"; were not aircraft size and the operational role impo-tant factors, and did not the mechanical limitations exert an undue influence? A number of speakers referred to auto-stabilization and the dangers of its indiscriminate use, and other points included the effect of p"oducts of inertia on the dynamic lateral stability, and the consideration of the complete aircraft rather than the wing alone in the study of aero-elastic effects. It was disclosed that the system of inter connected rudder and ailerons mentioned by the lecturer in connection with reducing adverse yaw at low speeds was in fact used on the Handley Page Victor. 1-0
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