Technical Memorandum X-26

.. .. 1 NASA TM X-26 TECHNICAL MEMORANDUM X-26 DESIGN GUIDE FOR PITCH-UP EVALUATION AND INVESTIGATION AT HIGH SUBSONIC SPEEDS OF POSSIBLE LlMITATIONS DUE TO WING-ASPECT-RATIO VARIATIONS By Kenneth P. Spreemann Langley Research Center Langley Field, Va. Declassified July 11, 1961 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON August 1959 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION . TEKX"NCAL MEMORANDUM x-26 DESIGN GUIDE FOR PITCH-UP EVALUATION AND INVESTIGATION AT HIGH SUBSONIC SPEEDS OF POSSIBLE LIMITATIONS DUE TO WING-ASPECT-RATIO VARIATIONS J By Kenneth P. Spreemann SUMMARY A design guide is suggested as a basis for indicating combinations of airplane design variables for which the possibilities of pitch-up are minimized for tail-behind-wing and tailless airplane configurations. The guide specifies wing plan forms that would be expected to show increased tail-off stability with increasing lift and plan forms that show decreased tail-off stability with increasing lift. Boundaries indicating tail-behind-wing positions that should be considered along with given tail-off characteristics also are suggested. An investigation of one possible limitation of the guide with respect to the effects of wing-aspect-ratio variations on the contribution to stability of a high tail has been made in the Langley high-speed 7- by 10-foot tunnel through a Mach number range from 0.60 to 0.92. The measured pitching-moment characteristics were found to be consistent with those of the design guide through the lift range for aspect ratios from 3.0 to 2.0. However, a configuration with an aspect ratio of 1.55 failed to provide the predicted pitch-up warning characterized by sharply increasing stability at the high lifts following the initial stall before pitching up. Thus, it appears that the design guide presented herein might not be applicable when the wing aspect ratio is lower than about 2.0. INTRODUCTION The phenomenon of longitudinal instability at high lift and the associated probability of an xndesirable divergence, commonly referred to as pitch-up, have presented a complex problem to designers of high- speed airplanes for many years. Generally, the designer is faced with the following considerations: (1) try to arrive at an aerodynamic lay- out for which pitch-up is unlikely or, at least, not severe, (2) evaluate the possible severity of the motion by means of a dynamic analysis of the type outlined in reference 1, and (3) if undesirable behavior is associated with the particular design that appears to be the best compromise of all requirements, provide sufficient artificial stabilization to mini- mize pitch-up as a problem. The present paper deals primarily with the first of these consider- L ations. The objectives are twofold: first, to present a design guide 4 (based on a considerable background of experience) to indicate combina- 2 tions of design variables for which the possibilities of pitch-up are 7 minimized; and second, to investigate one possible limitation of the guide with respect to the effect of wing-aspect-ratio variations on the tail contribution. The design guide presented is limited in applicability because the background information considered is predominately from inves- tigations concerned with fighter airplanes having wings with rather low aspect ratios and being of the tailless or tail-behind-wing types. The results, therefore, have their most direct application to such configurations having aspect ratios from about 2 to 5. COEFFICIENTS AND SWLS Figure 1 shows the stability system of axes employed for data pres- entation with arrows indicating positive directions of forces, moments, and angles. The coefficients and symbols used are defined as follows: cL lift coefficient, Lift/qS cD drag coefficient, Drag/qS c?rl pitching-moment coefficient, Pitching moment/qSE maximum lift coefficient 9 dynaqic pressure, pV2/2, lb/sq ft P rnass density of air, slugs/cu ft v free -s tream velocity , ft/sec I.I Mach number 3 S surface area, sq ft b span, ft C local chord parallel to plane of symmetry, ft - C mean aerodynamic chord, 2S /Ob/2 c2dy, ft L mean aerodynamic chord of vertical tail, ft 4 2 tail length (distance between quarter-chord points of wing and 7 horizontal-tail mean aerodynamic chords), ft X afterbody distance along body axis, ft spanwise distance from plane of symmetry on wing, ft tail position (height of tail above wing chord plane), ft diameter, in. r radius, in. A aspect ratio, b2/S h angle of sweep, deg %/4 angle of sweep of lifting-surface quarter chord, deg A taper ratio, Tip chord/Root chord R Reynolds number a angle of attack, deg i angle of incidence of fixed surface, deg E downwash angle induced by wing-fuselage combination, deg --- a, E% ac, c 4 Subscripts : max maximum t horizontal tail b model base Abbreviations : c.g. center-of-gravity location L 4 W wing 2 7 F fuselage V vertical tail H horizontal tail DESIGN GUIDE FOR PITCH-UP EVALUATION Wing and Wing-Fuselage Combinations The problem of pitching-moment nonlinearity at high lift coefficients for wings and wing-fuselage combinations has been studied by numerous investigators. For the most part, correlations have been made in terms of the wing sweep angle and aspect ratio, with varying degrees of consideration given to the taper ratio. Wing-thickness effects normally have been neglected, although the studies usually have been limited to wings that are thin enough to avoid any significant moment breaks due to thickness within the Mach number range investigated. Effects of cam- ber, twist, cranked plan forms, or any of the localized wing modifica- tions, such as fences, leading-edge extensions, or notches, have not been treated in any generalized sense. Such devices, though frequently very effective in improving pitching-moment linearity, normally must be tailored to the specific wing in question. A considerable number of such studies have been reported. It also has been comonly assumed that results on wings alone or on wing-fuselage combinations may be used inter- changeably, insofar as pitching-moment linearity is concerned. This appears to be justified for configurations having the proportions of manned aircraft in the subsonic and transonic speed ranges. An investigation (ref. 2) in which some rather extreme changes in forebody geometry were made showed little effect on the nature of the 5 moment breaks or on the angle of attack at which they occurred, but it indicated significant effects at angles of attack well beyond the initial moment breaks. For missile configurations, having large bodies in com- parison with the size of the wings, the fuselage characteristics may be predominant and, consequently, cannot be ignored. Some evidence exists which indicates that a tendency toward a similar situation may apply to supersonic-airplane configurations. (See ref. 3.) Several of the correlation boundaries that have been presented in the literature are summarized in figure 2. All are given in terms of coordinates of aspect ratio and sweep angle and in each instance the area above and to the right of the curve is characterized by decreased stability at the higher lift coefficients; whereas, the area below and to the left of the curve is characterized by increased stability at high lift. The range of data considered and the specific interpretation of each of the boundaries differ, however, as will be explained in the following paragraphs. The oldest and most widely known of the boundaries is that labeled @ " in figure 2 and given by Shortal and Maggin (ref. 4). Though based on only a limited amount of law-speed test data, it has served as a use- ful design criterion for many years. This boundary considers moment changes in the vicinity of maximum lift, without regard to slope changes in the intermediate lift range. Information obtained at a later date on more highly tapered wings showed that when attention was still con- fined to nonlinearities near maximum lift, a rather large effect of taper was indicated. Consequently, Furlong and McHugh (ref. 5) proposed I. an empirical expression for low-speed boundasies in which the taper ratio entered as a variable. The boundary evolved for A = 0 is indicated in figure 2 as the curve labeled @ .I1 No significant change from the Shortal-Maggin boundary was indicated for untapered wings. As wind-tunnel experience increased in the high subsonic and transonic speed range, the phenomenon of shock stall and the associated center-of-pressure changes were seen to impose limitations on the applicability of the boundaries derived from low-speed data. Studies made by Weil and Gray of the limited amount of high-speed data available in 1953 led to the proposal of a tentative boundary applicable to transonic speeds (ref. 6). This boundary, labeled "@" in figure 2, showed that in order to avoid stability reductions at the higher speeds, values of aspect ratio and sweep angle considerably lower than those indicated by the low-speed boundaries were required. The results available were insufficient to establish an effect of taper ratio, and the true signif- icance of the boundary as to the lift coefficients at which the moment 6 changes occurred was not clearly defined. Weil and Gray (ref. 6) pointed out the relation of their boundary to tail position. That is, wings represented by combinations of aspect ratio and sweep angle in the area above and to the right of the boundary probably could be used with a very low tail; whereas a very high tail may be feasible if wings represented by the area below and to the left of the boundary are used. This viewpoint inspired a systematic program aimed at defining the tail-off boundary more precisely and indicating fairly specific areas where the tail might be located.

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