1、NASA TECHNICAL NOTE EFFECTS OF WING LEADING-EDGE RADIUS AND REYNOLDS NUMBEK ON LONGITUDINAL AERODYNAMIC CHARACTERISTICS OF HIGHLY SWEPT WING-BODY CONFIGURATIONS AT SUBSONIC SPEEDS William P. Henderson Langley Research Center Hampton, Va. 23665 L nm NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHI
2、NGTON, D. C. DECEMBER 1976 B Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM I111111 11111 11111 11111llllllllllIIlH Ill1Ill1 -I. Report No. 2. Government Accession No. NASA TN D-8361 I i. Title and Subtitle EFFECTS OF WING LEAD
3、ING-EDGE RADIUS AND NUMBER ON LQNGITUDINAL AERODYNAMIC ISTICS OF HIGHLY SWEPT WING-BODY CONFIGURATIONS AT SUBSONIC SPEEDS 7. Author(s) 8. Performing Organization Report No. William P. Henderson I L- 11017 10. Work Unit No. 9. Performing Organization Name and Address 505-04-11-01 NASA Langley Researc
4、h Center 11. Contract or Grant No. Hampton, VA 23665 I 13. Type of Report and Period Covered 2. Sponsoring Agency Name and Address Technical Note National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, DC 20546 5. Supplementary Notes 6. Abstract An investigation was cond
5、ucted in the Langley low-turbulence pressure tunnel to deter mine the effects of wing leading-edge radius and Reynolds number on the longitudinal aerody namic characteristics of a series of highly swept wing-body configurations. The tests were conducted at Mach numbers below 0.30, angles of attack u
6、p to 16O, and Reynolds numbers per meter from 6.57 X lo6 to 43.27 X lo6. The wings under study in this investigation had leading-edge sweep angles of 61.7, 64.61, and 67.01 in combination with trailing-edge sweep angles of 0 and 40.6. The leading-edge radii of each wing planform could be varied from
7、 sharp to nearly round. 17. Key-Words (Suggested by Authoris) ) 18. Distribution Statement Reynolds number effects Unclassified - Unlimited Aerodynamics Leading-edge radius effects Subject Category 02 19. Security Classif. (of this r.?port) 1 21:. S+curity Classif. (of this page) I 21. NO. ; Pages I
8、 22. Price $4.25Unclassified 1Jncla ssified Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EFFECTS OF WING LEADING-EDGE RADIUS AND REYNOLDS NUMBER ON LONGITUDINAL AERODYNAMIC CHARACTERISTICS OF HIGHLY SWEPT WING-BODY CONFIGURATIONS AT SUBSONIC SPEED
9、S William P. Henderson Langley Research Center SUMMARY An investigation was conducted in the Langley low-turbulence pressure tunnel to determine the effects of wing leading-edge radius and Reynolds number on the longitudinal aerodynamic characteristics of a series of highly swept wing-body configura
10、tions. The tests were conducted at Mach numbers below 0.30, angles of attack up to,16, and Reynolds numbers per meter from 6.57 X 106 to 43.27 X lo6. The wings under study in this investi gation had leading-edge sweep angles of 61.7, 64.61, and 67.01 in combination with trailing-edge sweep angles of
11、 0 and 40.6. The leading-edge radii of each wing planform could be varied from sharp to nearly round. The results of this study indicate that for the sharp leading-edge wings, with a trailing-edge sweep angle of Oo, the experimental lift coefficient data are in excellent agreement with the theoretic
12、al estimates, potential flow plus leading-edge augmented vor tex lift, over the test angle-of-attack range. Changing the wing leading-edge shape from a sharp to a finite radius has a significant effect on the aerodynamic characteristics of the wing-fuselage configuration. Because of the development
13、of some leading-edge suction, The summationthe lift data lie between the potential and potential plus vortex estimates. of the experimental leading-edge suction and vortex lift, for wings of the same sweep angle but differing leading-edge shapes, is the same even though the individual increments are
14、 a function of leading-edge shape. As the leading-edge sweep or the trailing-edge sweep was increased for the wings with finite leading-edge radii, the summation of the experimental leading-edge suction and vortex lift was greater than the theoretical estimate. Increasing Reynolds number, leading-ed
15、ge radii, or trailing-edge sweep increased the angle of attack at which the experimental lift coefficient departed from the potential flow lift coefficient estimate. INTRODUCTION Aircraft of the future will probably be required to operate efficiently over a very large flight envelope. For example, a
16、ircraft designed for efficient supersonic cruise must Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-also be designed for efficient off-design performance. Off-design characteristics, such as take-off and landing requirements, subsonic cruise, and l
17、oiter characteristics, must be considered early in an aircraft design cycle so that the primary design goals can be achieved without adversely affecting the size or performance of the vehicle. The design principles associated with efficient supersonic design may not be compatible with either efficie
18、nt subsonic and transonic cruise or maneuvering characteristics. As technology in aircraft design has developed, methods for improving multimission capability have been sought. One such method, the subject of this paper, designs the wings to achieve fully attached (potential) flow at the cruise and
19、loiter conditions and controlled leading-edge separation (vortex flows) at maneuvering conditions. A significant amount of vortex lift is thereby achieved. There are, however, some indications that a wing designed to cruise with vortex lift present may be advantageous. With such a wing, the added li
20、ft results in a reduction in cruise angle of attack and, therefore, a reduction in cruise drag. Discus sions of the principle of vortex lift are presented in references 1and 2. Both attached (potential) flow at low angles of attack and full vortex flow at high angles of attack may be achieved by car
21、eful design of the wing leading-edge shape. This paper presents an analysis of data obtained for a series of wings (leading-edge sweep between 61.7 and 67.01) covering planforms of interest in the design of super sonic cruise vehicles. Each wing concept was studied with several leading-edge shapes o
22、ver a wide range of Reynolds numbers to determine the effect of these parameters on the longitudinal aerodynamic characteristics of the configuration. This study was conducted in the Langley low-turbulence pressure tunnel at Mach numbers less than 0.30 and angles of attack from -5 to 16. SYMBOLS The
23、 results given in this paper are referred to the stability axis system with the exception of the lift and drag coefficients, which are referred to the wind axis system. The force and moment data for each wing planform are nondimensionalized with respect to its own geometric characteristics (see fig.
24、 1). The moment reference center was located at a point 59.00 cm rearward of the nose along the model reference line (see fig. 1). cD drag coefficient, Drag 9- ref D,O drag at zero lift Lift CL lift coefficient, m r ef 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without li
25、cense from IHS-,-,-m pitching-moment coefficient, Pitching moment qmSref cN normal-force coefficient, Normal force qwSref cS leading-edge suction-force coefficient -mean geometric chord, cm kpotential potential-lift factor (see ref. 2) kv,le leading-edge vortex-lift factor (see ref. 2) s, free-strea
26、m dynamic pressure R Reynolds number, per meter Reynolds number based on wing mean geometric chord ref wing reference area, m 2 a angle of attack, deg OB angle at which lift coefficient departs from potential flow estimate, deg le wing leading-edge sweep angle, deg te wing trailing-edge sweep angle,
27、 deg MODEL DESCRIPTION Geometric details of the model are presented in figure 1. The configuration had a midwing with zero dihedral and a cylindrical fuselage with an ogive nose. The wing was composed of a slab-shaped center section to which various leading-edge and trailing-edge extensions could be
28、 attached. The leading-edge extensions, representing leading-edge sweep angles of 61.7, 64.61, and 67.01, were studied. For each leading-edge sweep, 3 C Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-a maximum of three leading-edge shapes were avail
29、able for study: one shape had a sharp wedge section, and the other two shapes had rounded sections with average leading-edge radii (perpendicular to the wing leading edge) of 0.050 cm and 0.121 cm, respectively. For each leading-edge configuration, two trailing-edge extensions were available: one wi
30、th a trailing-edge sweep angle of Oo, the second with a trailing-edge sweep angle of 40.6. TESTS AND CORRECTIONS Tests were conducted in the Langley low-turbulence pressure tunnel at Mach num bers below 0.30 and at angles of attack of up to 16. The test Reynolds number per meter varied from 6.57 X l
31、o6 to 43.27 X lo6. Transition strips, 0.32 cm wide, of No. 150 carbo rundum grains were placed 1.00 cm behind the leading edge of the wings and 2.54 cm behind the nose of the fuselage. Corrections to the model angle of attack were made for deflections of the balance and sting support system under ae
32、rodynamic load. The lift coef ficient and drag coefficient data were corrected for jet boundary and blockage effect. The drag data were adjusted to correspond to free-stream static conditions in the balance chamber. No attempt was made to correct data for any possible aeroelastic distortion caused b
33、y load at high dynamic pressures. PRESENTATION OF RESULTS The basic data for each configuration studied are presented in figures 2 to 13 and are summarized in figures 14 to 18. As an aid in locating a particular part of the data, the following index of figures is presented. Figure Effect of Reynolds
34、 number on longitudinal aerodynamic characteristics of configuration with -Ale = 61.7O, Ate = Oo, and sharp leading edge . Ale = 61.7, Ate = Oo, and small leading-edge radius . Ale = 61.7, Ate = Oo, and large leading-edge radius . Ale = 61.7, Ate = 40.6, and sharp leading edge . Ale = 61.7, Ate = 40
35、.6, and large leading-edge radius . Ale = 64.61, Ate = Oo, and sharp leading edge Ale = 64.61, Ate = Oo, and large leading-edge radius Ale = 67.01, Ate = Oo, and sharp leading edge Ale = 67.01, Ate = Oo, and small leading-edge radius 2 3 4 5 6 7 8 9 10 4 Provided by IHSNot for ResaleNo reproduction
36、or networking permitted without license from IHS-,-,-= 67.01, Ate = Oo, and large leading-edge radius = 67.Oloy Ate = 40.6, and sharp leading edge = 67.01, Ate = 40.6, and large leading-edge radius Effect of leading-edge profile on leading-edge vortex flow characteristics for configuration with lead
37、ing-edge sweep of 61.7 and trailing-edge sweepof0O. Effect of leading-edge sweep on leading-edge vortex flow characteristics Figure 11 12 13 14 15 16 17 18 for configuration with trailing-edge sweep of 0 . Effect of trailing-edge sweep on leading-edge vortex flow characteristics for configuration wi
38、th leading-edge sweep of 61.7 and large leading-edge radius Effect of Reynolds number on leading-edge vortex flow characteristics for configuration with leading-edge sweep of 61.7, trailing-edge sweep of Oo, and large leading-edge radius . Effect of leading-edge radius and trailing-edge sweep on var
39、iation of angle of attack at which leading-edge separation starts as function of Reynolds number RESULTS AND DISCUSSION Since the drag data presented here have been previously analyzed and presented in the published literature (see ref. 3), only the general trends of the data are discussed in this p
40、aper, and the primary discussion considers only the lift and pitching-moment coeffi cient data. The lift and pitching-moment data on each figure are compared primarily with two theoretical estimates; one estimate is based on potential flow, and the other estimate is based on potential plus vortex fl
41、ow. The potential flow estimate was made by using the method of reference 4, and the vortex flow estimate was made by using the method of ref erence 2. The estimated lift coefficients for potential flow presented in these figures do not include the contribution of the leading-edge suction to the lif
42、t. To illustrate the effect of this component on the lift, the leading-edge suction contribution to lift calculated by $,le cos Ale sin3 a, is plotted in figures 3 and 11 at the lowest Reynolds number. As expected, these data indicated that the contribution of leading-edge suction to lift is very sm
43、all for the wings of 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-this study at the test angle-of-attack range. The vortex-lift increment is the contribution of the leading edge and the augmented portions. The concept of augmented vortex lift wa
44、s developed because for many delta Wings the leading-edge vortex generated on the wing per sists for a considerable distance downstream and, therefore, can act on other surfaces, such as the aft part of wing planforms. No contribution of the side edge was included because of the high taper ratio of
45、the Wing, which naturally results in very small wing tip chord for any direct generation of vortex lift. The experimental data are compared with the theoretical estimates in a limited number of cases to reduce the clutter in the figures. The drag data for the wings with the sharp leading edges (see
46、figs. 2, 5, 7, 9,and 12) agree very well with the theory for potential flow plus vortex flow over the entire test lift coefficient range investigated. These characteristics should be expected since sharp leading-edge wings do not develop any leading-edge suction. As either the leading-edge radius or
47、 Reynolds number is increased for wings with finite leading-edge radii, leading-edge suction is developed, and the data agree with the potential flow theory over a particu lar lift coefficient range which varies. As indicated in reference 3 and illustrated in these data (see fig. 8, for example), th
48、e lift coefficient at which the drag data depart from the potential flow theory is highly dependent on both Reynolds number and leading-edge radius. For all three sharp leading-edge wings with a trailing-edge sweep angle of 0 (see figs. 2, 7, and 9), the agreement of the experimental lift data with
49、the potential flow plus vortex flow theory is excellent over the entire test angle-of-attack range. In this application of the vortex theory, only the leading edge and augmentation effects are accounted for. There fore, the relative insignificance of the side effects for most of these highly tapered wings seems to be substantiated. Th
