NASA-TN-D-2824-1965 Large-scale wind-tunnel investigation of the low-speed aerodynamic characteristics of a supersonic transport model having variable-sweep wings《带有可变掠翼的超音速运输机模型低速.pdf

1、NASA TECHNICAL NOTE TN D-2824 LC_ LARGE-SCALE WIND-TUNNEL INVESTIGATION OF THE LOW-SPEED AERODYNAMIC CHARACTERISTICS OF A SUPERSONIC TRANSPORT MODEL HAVING VARIABLE-SWEEP WINGS by Anthony M. Cook, Ames Reseurcb Center Moffett Field, Cui$ Richurd K. Gre$ und Kijoshi Aoyugi - NATIONAL AERONAUTICS AND

2、SPACE ADMINISTRATION WASHINGTOW -_ MAY 1965 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NASA TN D-2824 TECH LIBRARY KAFB, NM I llll1l1llll I lllll Ill11 lllll1llll Ill1 Ill1 0079b48 LARGE -SCALE WIND -TUNNEL INVESTIGATION OF THE LOW-SPEED AERODYN

3、AMIC CHARACTERISTICS OF A SUPERSONIC TRANSPORT MODEL HAVING VARIABLE-SWEEP WINGS By Anthony M. Cook, Richard K. Greif, and Kiyoshi Aoyagi Ames Research Center Moffett Field, Calif. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Scientific and Technical Inform

4、ation Springfield, Virginia 22151 - Price $4.00 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-LARGE-SCALE WIND-TUNNEL INVESTIGATION OF TRE LOW-SPEED AERODYNAMIC CHARACTERISTICS OF A SUPERSONIC TRANSPORT MODEL HAVING VARIABLELSWEEP WINGS* By Anthony

5、 M. Cook, Richard K. Greif, and Kiyoshi Aoyagi Ames Research Center SUMMARY The results are presented as six-component aerodynamic force and moment data obtained at various angles of attack and sideslip. a Reynolds number of 16 million, based upon the mean aerodynamic chord of the wing swept to 750.

6、 and aspect ratio, leading-edge slat deflection and geometry, trailing-edge flap deflection, geometry, and span extent, and horizontal-tail geometry. Data were obtained at The investigation included variations of wing sweepback The results show that all configurations tested, except one, were longi-

7、 The configuration that was not unstable had tudinally unstable at high lift. a tail in a low horizontal position, a wing sweepback angle of 23 with a large portion of the fixed wing deflected as a leading-edge flap. INTRODUCTION The development of any supersonic aircraft involves combining aerodyna

8、mi- cally incompatible high- and low-speed design requirements. sweep wing concept is one approach to this problem. One basic requirement in this approach is to provide acceptable stability characteristics by minimizing the aerodynamic center shift due to wing sweep. The variable- Earlier concepts o

9、f variable-sweep wings (ref. 1) incorporated a longi- tudinal translation of the wing together with change in sweep angle to elimi- nate the aerodynamic center shift associated with changing sweep. Efforts to avoid the mechanical difficulties inherent with longitudinal translation of the wing result

10、ed in the concept of the fixed outboard pivot and a fixed, highly swept, inboard wing section designed to minimize aerodynamic center shift (refs. 2 through 6). instability characteristics at the stall for the high-lift configurations of this design. longitudinal instability and the maximum lift cha

11、racteristics of high-lift, variable-sweep configurations at high Reynolds numbers. Small-scale results give evidence of longitudinal The purpose of the tests reported herein was to investigate this The scope of this investigation was limited to the first-order effects of the variables considered mos

12、t important: wing sweep in low-speed cruise and -. - . - i - *itle, Unclassified. I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-high-lift configurations, wing aspect ratio, trailing-edge flap systems, leading-edge slats, horizontal-tail area and

13、location, and fixed-wing leading- edge radius and flaps. NOTATION A AR ac b CD CL Cl Cm Cn C - C FDS FZE FSS gS iT L/D LE wing area (see Reduction of Data), sq ft b2 aspect ratio, - A aerodynamic center wing span, ft drag coefficient, drag SA lift lift coefficient, - SA rolling moment rolling-moment

14、 coefficient, qAb pitching moment qAE pitching-moment coefficient , awing moment yawing-moment coefficient, Y 9Ab side force SA side-force coefficient, b/2 chord n mean aerodynamic chord, eJ c2 dy, ft flap, double slotted fixed-wing leading-edge flap flap, single slotted gap of leading-edge slats, f

15、raction of chord horizontal-tail incidence (positive when trailing edge is down), deg lift-drag ratio leading edge 2 . _ i_ _._ .-. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2T rFZE VT - X Y Z a P 6 av 11 A *FL and two low-aspect-ratio configur

16、ations, with high and mid horizontal-tail positions (configurations Various wing leading-edge sweepback angles ranging from l3-l/Z0 two high-aspect-ratio configurations, Four configurations are shown installed in the wind A1 and A2, BL and B2) . The wing pivot was located at 36-percent semispan and

17、46-percent chord of the fully swept wing (based upon the low-aspect-ratio wing of configura- tion B). The fixed portion of the wing was provided with either TO0 or 750 leading-edge sweep. Planform Geometry Geometric details of the high-aspect-ratio configuration (A) and the low- aspect-ratio configu

18、ration (B) can be found in tables I and 11, respectively. A sketch including pertinent dimensions of the model is shown in figure 2. The airfoil section for the movable wing had a flat lower surface and the thickness distribution of an NACA 65006 airfoil section. for wing airfoil coordinates. See ta

19、ble 111 The lower aspect ratio of configuration B was obtained by removing 3-1/2 feet of wing tip from configuration A. Fixed-wing section geometry is detailed in figure 2(c) by cross sections at various fuselage stations. The basic leading edge was sharp along its entire length. However, an alterna

20、te, rounded leading edge shown in figure 2(e) was also tested. This rounded leading edge tapered from a radius of 3 inches at the fuselage junc- ture to 0.73 inch (wing leading-edge radius at movable-wing juncture). Planform details are given in figure 2(d). The fuselage consisted of a blended wing-

21、body section, as shown in figure 2(c), with an underslung, side-by-side engine nacelle with plugged, two-dimensional inlets faired to the rectangular aft fuselage shown in figure 1. Horizontal Tail The horizontal tail was tested in three positions (see fig. 2(b): low, In the low position it was moun

22、ted on the fuselage at 10 per- C (of 25 sweep) below the wing-chord plane; in the mid position it was mid, and high. cent mounted on the vertical stabilizer at LO percent C above the wing-chord plane; in the high position it was also mounted on the vertical stabilizer, 4 Provided by IHSNot for Resal

23、eNo reproduction or networking permitted without license from IHS-,-,- at 50 percent c above the wing-chord plane. Because of the sweepback of the vertical stabilizer, horizontal-tail length ( 2) varied for the three posi- tions. Two horizontal-tail sizes were tested in the high position. For all te

24、sts of configuration Al, the low tail was at a negative dihedral of 100. High-Lift Devices Fixed-wing high-lift devices.- Details of the plain flap of the fixed wing are shown in figure 2(d). A simulated fiker type flap was tested on the leading edge of the fixed wing, with both sharp and rounded fi

25、xed-wing leading edge (see fig. 2(e). Movable-wing trailing-edge double-slotted flap system.- The double- slotted flap geometry and-a typical -cross section are shown in figure 2(f). The vane was 7-l/2 percent of the wing chord, streamwise, with the wing at 25 sweep. 2-percent wing chord was maintai

26、ned at the vane. Flap deflections ranged from 30 to 600 in loo increments. flap performance. The modification (fig. 2(f) consisted of adding sheet metal extensions to the wing trailing-edge shroud aqd vane and was used for all tests of double-slotted flaps unless otherwise noted. The main flap compr

27、ised 25 percent of the wing chord. A slot of The slot geometry was modified to improve Movable-wing trailing-edge . - . single-slotted flap system.- The single- slotted flap configuration was achieved by removal of the vane of the double- slotted flap and moving the flap forward into the wing. wing

28、chord by 4 percent and accounts for the difference in wing area and aspect ratio between the two flap systems. A slot of 2-percent wing chord was maintained at all flap deflections, and the range of flap deflection was from 0 to 300, bo0, and 50. The geometry and cross-section details of this flap s

29、ystem are given in figure 2(g). This reduced the Both flap systems were constructed in three sections, extending (as shown in fig. 2(a) from 20 to 52 percent semispan, from 52 to 67 percent semispan, and from 67 to 98 percent semispan of the high-aspect-ratio wing. result, flap deflection notation i

30、s indicated in three parts: As a 6 = inboard deflection/middle deflect ion/outboard deflect ion Movable-wing leading-edge slats.- The details of leading-edge slat size, deflection, and positioning are shown in figure 2(h). Two sized slats were tested, one having a length equal to l5-percent streamwi

31、se wing chord (at 25 sweep), and the other, 18-3/4-percent wing chord. The profile of the 0.15 slat was made to match the leading-edge profile of the wing. The 0.1875 slat incorporates the basic 0.15 slat with a rounded leading-edge extension to provide camber as shown in the figure. Slat deflection

32、, 6s1 is given relative to its undeflected position as if it were “gloved“ onto the 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-wing. Slat gap, gs, was varied from 0- to 2-percent chord in 1/2-percent increments. Unless otherwise noted, all sla

33、t data reflect the use of the basic slat of 0.15 length. TESTING AND PROCEDURE Six-component force and moment data were obtained by conventional wind- tunnel testing methods through an angle-of-attack range from -4 to +22O, and an angle of sideslip from -2 to A0. 15 pounds per square foot, correspon

34、ding to a Reynolds number of 16 million, based upon mean aerodynamic chord at 75O wing sweep. Free-stream dynamic pressure was The majority of tests were directed toward the development of high-lift devices and the investigation of longitudinal stability characteristics for landing and take-off conf

35、igurations. REDUCTION OF DATA Corrections Standard corrections were applied to angle of attack to account for wind- tunnel wall effects. The corrections accounted for the variations in span due to wing sweep. Measured drag was corrected in accordance with the angle-of- attack correction. measurement

36、s to account for strut tares: In addition, the following correction was added to drag No correction was mde for tunnel-wall corrections for tail-on conditions due to the variable-sweep nature of the configuration. Reference Dimensions The computation of force and moment coefficients for all wing swe

37、eps of a given configuration was based on the dimensions corresponding to the total wing area, including fixed wing, at the 75O sweep condition of that particular configuration . Moment Center The moment center for all configurations, regardless of wing sweep, was taken on the axis of the wing pivot

38、, 2.875 inches above the wing-chord plane. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-RESULTS The acquisition of data for this investigation covered four testing periods and the several configurations previously mentioned. The results in figures

39、 3 through 48 present longitudinal characteristics and selected cases include lateral-directional characteristics. These results are swmnarized in figures 49 through 56 and are discussed more completely in the Discussion section of this report. Table IV is a complete index to the figures. DISCUSS IO

40、N General Characteristics Aspect ratio.- Figure 49 presents a comparison of the test data for wings of aspect ratio 6.9 and 8.4 at 25O of wing leading-edge sweepback and with high-lift devices installed. At 12O angle of attack, for instance, there is an incremental loss of 6.5 percent CL in reducing

41、 aspect ratio from 8.4 to 6.9, accompanied by a decrease in stability of -percent static margin. The reduction in aspect ratio causes a lower lift-curve slope, but it is shown that for this degree of wing sweepback, there is essentially no difference in ChX * Wing sweep.- Figure 50 shows the effects

42、 of 13-1/2O and 25O wing sweep for both the flaps-up and flaps-down conditions. deflection, it is seen that there is no appreciable benefit to be derived by a wing sweepback angle of less than 25O, in terms of a “usable“ that CL at which pitch-up occurs. Changes in lift due to wing sweep for the fla

43、ps-up condition are also very small. Note that the aerodynamic-center shift due to wing sweep from l3-l/z0 to 25 with flaps up amounts to 8-percent static margin and is essentially the same as the change in static margin du.e to 40 of flap deflection at l3-l/Z0 of wing sweep. (The static margin chan

44、ge d.x tG flap; deflection, however, is a result of the downwash flow at the particular horizontal-tail location, since no change is indicated in the tail- off d-a.t.a of figs. 36 md 37. ) In the case of 400 flap Cbx, or Longitudinal Stability As mentioned in the Introduction, variable-sweep configu

45、rations generally haTvre unstable pitching-moment characteristics at high lift coefficients. The reason is that a wing-tip stall progressing inboard (based on tuft observa- tions) is further a.ggra.w,ted by a vortex generated along the highly swept leading edge Of tne fixed wing delaying idoard stal

46、l. The size and sweep of the fixed-wing portion contribute to the strength of this vortex. Part of this imestigation involve3 testing horizcztzl -tzil pscitionc in coziaination wih baiious Cluw conLro1 Gevices in vr Becht, Robert E.; and Few, Albert G., Jr.: Stability and Control Characteristics at

47、Low Speed of a l/kScale Bell X-5 Airplane Model. Longitudinal Stability and Control. NACA RM 908, 1950 2. Alford, William J., Jr.; and Henderson, William P.: An Exploratory Investigation of the Low-Speed Aerodynamic Characteristics of Variable- Wing-Sweep Airplane Configurations. NASA TM X-142, 1959

48、. 3. Spencer, Bernard, Jr.: Stability and Control Characteristics at Low Subsonic Speeds of an Airplane Configuration Having Two Types of Variable-Sweep Wings. NASA TM X-303, 1960. 4. Spencer, Bernard, Jr .: Low-Speed Longitudinal Aerodynamic Characteristics Associated With Variations in the Geometr

49、y of the Fixed Portion of a Variable-Wing-Sweep Airplane Configuration Having an Outboard Pivot. NASA TM x-625, 1962. 3. Alford, Williazz VT., ZF.; Luora, Avro A.; and Hendersoii, William P.: Wind-Tunnel Studies at Subsonic and Transonic Speeds of a Multiple- Mission Variable-Wing-Sweep Airplsce CGrfigT2ation. NASA TM x-206, 1959 6. Foster, Gerald V.; a