1、I NASA TECHNICAL NOTE NASA TN D-7099 I l o* 0 h 1- z c I EFFECT OF WING DESIGN ON THE LONGITUDINAL AERODYNAMIC CHARACTERISTICS OF A WING-BODY MODEL AT SUBSONIC SPEEDS by Wikm P. Henderson und Jmrett K, Hzifffi2un Luqley Research Center Hamptotz, Vu. 23365 NATIONAL AERONAUTICS AND SPACE ADMINISTRATIO
2、N WASHINGTON, D. C. DECEMBER 1972 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1. RepOrt No. 2. Government Accession No. NASA TN D-7099 4. Title and Subtitle December 1972 EFFECT OF WING DESIGN ON THE LONGITUDINAL MODEL AT SUBSONIC SPEEDS AERODYNA
3、MIC CHARACTERISTICS OF A WING-BODY 3. Recipients Catalog No. 5. Rewrt Date 7. Author(s) William P. Henderson and Jarrett K. Huffman 9. Performing Organization Name and Address NASA Langley Research Center Hampton, Va. 23365 8. Performing Organization Report No. L-8562 10. Work Unit No. 760-67-01-01
4、11. Contract or Grant No. 16. Abstract 12. Sponsoring Agency Name and Address National Aeronautics and Space Administration Washington, D.C. 20546 An investigation has been conducted to determine the effects of wing camber and twist on the longitudinal aerodynamic characteristics of a wing-body conf
5、iguration. Three wings were used, each having the same planform (aspect ratio of 2.5 and leading-edge sweep angle of 44) but differing in amounts of camber and twist (wing design lift coefficient). The wing design lift coefficients were 0, 0.35, and 0.70. over a Mach number range from 0.20 to 0.70 a
6、t angles of attack up to about 22. The effect of wing strakes on the aerodynamic characteristics of the cambered wings was also studied. A comparison of the experimentally determined aerodynamic characteristics with theoretical estimates is also included. The investigation was conducted 13. Type of
7、Report and Period Covered Technical Note 14. Sponsoring Agency Code 17. Key Words (Suggested by Author(s) Cambered wings Longitudinal aerodynamic characteristics . thus, the tendency for flow separation at the design lift coefficients is suppressed. This concept was investigated for a range of desig
8、n lift coefficients up to 0.7. The use of leading- and trailing-edge flaps to approximate this concept was investigated in reference 1. , I The second concept makes use of the vortex lift produced by leading-edge separa- I tion from sharp highly swept wing strakes. The success of this concept depend
9、s on the I I mutual interaction of the strake vortex and the main wing which is difficult to predict analytically. These two concepts were investigated individually and in combination in subsonic wind-tunnel tests using a wing-body model. Experimental lift and drag coef- ficient characteristics obta
10、ined with the wing strake and an uncambered wing are com- pared with analytical predictions based on the leading-edge-suction analogy of vortex lift. I This study was conducted in the Langley high-speed 7- by 10-foot tunnel at Mach I numbers from 0.20 to 0.70 and at angles of attack up to 22. SYMBOL
11、S I The results as presented are referred to the body axis system with the exception of the lift and drag coefficients, which are referred to the wind axis system. The moment reference center was located at a point 65.91 Centimeters rearward of the nose (long fuselage) along the model reference line
12、s. I I (See fig. 1.) A aspect ratio b wing span, centimeters I cD drag coefficient, Drag qs ACD increment in drag associated with addition of wing strake drag coefficient at zero lift D,o 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-cL ACL L,d C
13、m C - C M n R S S X Y z 01 “i r Lift lift coefficient, - qs increment in lift associated with addition of wing strake wing design lift coefficient Pitching moment pitching-moment coefficient, qsc local wing chord, centimeters wing mean geometric chord, 23.30 centimeters Mach number nth loading funct
14、ion free-stream dynamic pressure, newtons per meter 2 Reynolds number (based on E) wing reference area, 1.0322 meters2 CL tan - (cD - CD o) 7 rai d leading -edge -suction parameter , CL tan Q! - A b/2 distance behind leading edge of wing, centimeters distance from fuselage reference line (measured s
15、panwise), centimeters wing airfoil ordinate, centimeters angle of attack, degrees induced angle of attack circulation strength 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 angle used to locate pressure doublets chordwise, 0 at leading edge and
16、 7 at trailing edge . Sub scripts: 1 lower surface U upper surface MODEL DESCRIPTION A three-view drawing of the basic model is presented in figure l(a) and a drawing showing the model with the wing strake is presented in figure l(b). A photograph of the model sting mounted in the Langley high-speed
17、 7- by 10-foot tunnel is presented in fig- ure 2. The model as illustrated in figure l(a) consists of a simple wing-fuselage config- uration, with the wing having an aspect ratio of 2.5, a taper ratio of 0.20, a wing leading- edge sweep angle of 44O, and NACA 64A series airfoil sections (measured st
18、reamwise) with a thickness ratio of 6 percent at the fuselage juncture and 4 percent at the wing tip. Three variations in wing camber and twist, corresponding to design lift coefficients of 0, 0.35, and 0.70 were studied. Ordinates for the cambered airfoils are presented in table I. Two fuselage len
19、gths were studied; the long fuselage was 11.94 cm longer than the short fuselage. The wing strake (fig. l(b) was constructed of a 0.159-cm-thick flat plate with sharp leading edges. The sharp leading edge had a total bevel angle of 3.2. WING DESIGN PROCEDURE The mean camber surfaces of the two cambe
20、red and twisted wings were designed by using the procedure of reference 2 for design points corresponding to CL of 0.35 or 0.70 at a Mach number of 0.40. At the design points, an elliptical span-load distribu- tion was specified and the chordwise load distribution was specified as the superposition
21、of four sin ne pressure modes with n = 1, 3, 5, and 7, where 2x COS e= The magnitude of the modes at each spanwise station was selected to approximate a rectangular chordwise load distribution. The resulting distribution is characterized by zero load at the leading edge and a very small region of st
22、rong adverse pressure gra- dient in the vicinity of the trailing edge. No camber was incorporated in the fuselage. 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TEST AND CORRECTIONS The investigation was conducted in the Langley high-speed 7- by
23、10-foot tunnel at Mach numbers from 0.20 to 0.70 and at angles of attack up to 22. The variation of the test Reynolds number, based on the wing mean geometric chord, with Mach number is presented in figure 3. Transition strips 0.32 cm wide of No. 100 carborundum grains (based on analysis of ref. 3)
24、were placed 1.14 cm streamwise from the leading edge of the wings and 2.54 cm behind the nose of the fuselage. Corrections to the model angle of attack have been made for deflections of the balance and sting support system due to aerodynamic load. Pressure measurements obtained from orifices located
25、 within the fuselage base cavity were used to adjust the drag coefficient to a condition of free-stream static pressure at the model base. Jet-boundary and blockage corrections estimated by the procedures of references 4 and 5, respectively, were applied to the data. PRESENTATION OF RESULTS Effect o
26、f wing design lift coefficient on aerodynamic characteristics of Effect of wing design lift coefficient on drag due to lift characteristics Combined effect of Mach number and Reynolds number on aerodynamic Effect of fuselage forebody length on aerodynamic characteristics of model Effect of wing stra
27、ke on aerodynamic characteristics of model with long model with short fuselage forebody and wing strake off of model with short fuselage forebody and wing strake off characteristics of model with short fuselage forebody and wing strake off with wing strake off. CL d = 0.70 fuselage forebody. CL,d =
28、0; M = 0.40 . 9 Effect of wing strake on aerodynamic characteristics of model. CL,d = 0.35 . Effect of wing strake on aerodynamic characteristics of model with long Increment in lift and drag due to adding wing strake. M = 0.40 Comparison of experimental and estimated lift and drag for model with fu
29、selage forebody. CL,d = 0.70 strake. CL,d=O . The basic longitudinal aerodynamic characteristics are presented in figures 4 to 10 and pertinent results are summarized in figures 11 and 12. As an aid in locating a par- ticular part of the data, the following list of figures is presented: Figure 4 5 6
30、 7 8 9 10 11 12 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-RESULTS AND DISCUSSION The effect of wing camber and twist on the longitudinal aerodynamic character- istics of the wing-body configuration at Mach numbers from 0.20 to 0.70 is present
31、ed in figure 4. An analysis of the effect of camber on the drag due to lift is presented in fig- ure 5 and the combined effects of Mach number and Reynolds number on the aerodynamic characteristics of the model are presented in figure 6. Increasing the wing camber and twist (CL d) results in a signi
32、ficant increase in the maximum lift coefficient obtained, with essentially no change in lift-curve slope between the uncambered and moderately cambered wing (CL,d = 0.35). The highly cambered wing (CL d = 0.70) showed some non- linearity in lift-curve slope over most of the angle-of-attack range; th
33、us, the presence of flow separation on the wing was indicated. Increasing the wing camber and twist (CL,d) is shown in figure 4 to result in rela- tively large values of nose-down pitching moment with some slight change in the pitching- moment-coefficient curve slope in the low lift coefficient rang
34、e. could result in trim drag problems for an airplane configuration, unless careful attention is given to the location of the center of gravity of the airplane and to the size of the longi- tudinal control surface. The nose-down pitch At the higher lift coefficients the uncambered wing (CL d = 0) sh
35、ows a severe increase in stability (pitch down) which is eliminated as the wing design lift coefficient is increased. The drag characteristics illustrated in figure 4 stow that cambering the wing, as would be expected, results in a significant increase in the drag at zero lift coefficient and a cons
36、iderable reduction in drag at the higher lift coefficients. analysis of the effects of camber on the drag characteristics, in that the drag of each wing is compared with the zero and full leading-edge-suction boundaries (lower part of fig. 5). The percent leading-edge suction developed by each wing
37、is also presented for comparison purposes in this figure. The two theoretical boundaries and the equation for computing leading-edge suction are discussed in reference 6. The leading-edge- suction parameter s as presented here simply represents the location of the experi- mental drag data relative t
38、o the two theoretical boundaries. For simplicity, only the zero leading-edge-suction boundary for the uncambered wing is presented in figure 5. The zero leading-edge-suction boundaries for the cambered wings would be essentially the same in the low lift range as that presented in figure 5. At the hi
39、gher lift coeffi- cients, the zero leading-edge-suction boundaries for the cambered wings would be slightly lower because of the improved lift-curve slope exhibited by the cambered wings. The data obtained on the uncambered wing model, because of extensive leading-edge sep- aration, depart froni the
40、 full suction boundary at a relatively low lift coefficient. The Figure 5 presents an 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-cambered wings, however, maintain nearly full leading-edge suction over a fairly wide lift coefficient range. exhi
41、bits higher than 90 percent suction over a range of lift coefficients from about 0.25 to 0.60 (fig. 5(b), whereas the wing with a design lift coefficient of 0.70 maintains over 90 percent suction up to a lift coefficient of nearly 0.90. For example, the wing with a design lift coefficient of 0.35 Th
42、e combined effects of Mach number and Reynolds number on the aerodynamic characteristics for the cambered wings of this study are presented in figure 6. Fig- ure 3 presents the variation of Reynolds number with Mach number. The effects of Mach number and Reynolds number cannot be determined independ
43、ently during these tests since the Langley 7- by 10-foot high-speed tunnel is an atmospheric tunnel and variations in Mach number are accompanied by changes in Reynolds number. The uncambered wing (fig. 6) shows a slight change in the longitudinal aerodynamic characteristics with increasing Mach num
44、ber and Reynolds number. The most significant effect appears to be in the decrease in the maximum lift coefficient (fig. 6(a) and the high angle-of-attack nose -down pitching -moment characteristics (fig. 6(c). As the design lift is increased (increased camber and twist), the effects of Mach number
45、and Reynolds number are considerably more pronounced. For example, increasing the Mach number and Reynolds number on the highly cambered wing (CL,d = 0.70) evidently reduces the extensive flow separation on the wing which, as shown in figure 6, results in a large increase in lift, a decrease in drag
46、, and a nose-down pitching-moment increment. The effect of fuselage forebody length on the aerodynamic characteristics of the model is illustrated by the data presented in figure 7. Lengthening the forebody, which was necessary to accommodate the wing strake (fig. l(b), increased the drag slightly (
47、increased wetted area) and reduced the stability level by about 2 percent 6. The increased wetted area accounted for more than half the drag obtained for the longer fuse lage forebody. The effect of forebody length is presented only for the highly cambered wing (CL,d = 0.70). It is believed, however
48、, that the effect illustrated for the highly cambered wing would be essentially the same for the other wings of this study. The effect of adding a wing strake in combination with the cambered wings on the aerodynamic characteristics of the model is presented in figures 8 to 10 and summarized in figu
49、re 11. Adding the wing strake (fig. 8) has essentially no effect on the lift coeffi- cient in the low angle-of-attack range (up to about 5). The increase in lift coefficient expected as a. resilt nf the added area of the wing strake is probably compensated for by a detrimental interference effect of the flow field generated by the strake on the wing. At the higher angles of attack adding the wing strake re
