NASA-TN-D-6263-1971 Lift and drag characteristics of the HL-10 lifting body during subsonic gliding flight《HL-10提升机身在亚音速滑翔飞行时的升力和阻力特性》.pdf

《NASA-TN-D-6263-1971 Lift and drag characteristics of the HL-10 lifting body during subsonic gliding flight《HL-10提升机身在亚音速滑翔飞行时的升力和阻力特性》.pdf》由会员分享，可在线阅读，更多相关《NASA-TN-D-6263-1971 Lift and drag characteristics of the HL-10 lifting body during subsonic gliding flight《HL-10提升机身在亚音速滑翔飞行时的升力和阻力特性》.pdf（28页珍藏版）》请在麦多课文档分享上搜索。

1、NASA TECHNICAL NOTE NASA TN D-6263- _- e,.i KIRTLAkD AFB,N.M, LIFT AND DRAG CHARACTERISTICS . OF THE HL-IO LIFTING BODY DURING SUBSONIC GLIDING FLIGHT b . hby Jon S. Pyle Flight Research Center Edwards, Cali$ 93523 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, 0. C. MARCH 1971 Provided b

2、y IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1. Report No. 2. Government Accession No. NASA TN D-6263 I 4. Title and Subtitle LIFT AND DRAG CHARACTERISTICS OF THE HL-10 LIFTING BODY DURING SUBSONIC GLIDING FLIGHT 7. Author(s) Jon S. Pyle ?-9. Performing Org

3、anization Name and Address NASA Flight Research Center P. 0. Box 273 Edwards, California 93523 d 112. Sponsoring Agency Name and Address National Aeronautics and Space Administration Washington, D. C. 20546 15. Supplementary Notes 16. Abstract TECH LIBRARY KAFB, NM I111111 11111 11111 11111 /Ill/lll

4、ll1111111111111 0333074 3. Recipients Catalog No. 5. Report Date March 1973 6. Performing Organization Code 8. Performing Organization Report No. H-608 10. Work Unit No. 727-00-00-01-24 11. Contract or Grant No. 13. Type of Report and Period Covered Technical Note 14. Sponsoring Agency Code Subsonic

5、 lift and drag data obtained during the HL-10 lifting body glide flight program are presented for four configurations for angles of attack from 5“ to 26“ and Mach numbers from 0.35 to 0. 62. These flight data, where applicable, are compared with results from small-scale wind-tunnel tests of an HL-10

6、 model, full-scale wind-tunnel results obtained with the flight vehicle, and flight results for the M2-F2 lifting body. The lift and drag characteristics obtained from the HL-10 flight results showed that a severe flow problem existed on the upper surface of the vehicle during the first flight test.

7、 This problem was corrected by modifying the leading edges of the tip fins. The vehicle attained lift-drag ratios as high as 4.0 during the landing flare (performed with the landing gear up), which is approximately 14 percent higher than demonstrated by the M2-F2 vehicle in similar maneuvers. 17. Ke

8、y Words (Suggested by Author(s) 18. Distribution Statement Lifting body - Lift-drag ratio - HL-10 vehicle -Flow separation Unclassified - Unlimited 19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. NO. of Pages 22. Price* Unclassified I Unclassified 25 I $3.00 For Sale

9、by the National Technical Information Service, Springfield, Virginia 22151 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-LIFT AND DRAG CHARACTERISTICS OF THE HL-10 LIFTING BODY DURING SUBSONIC GLIDING FLIGHT Jon S. Pyle Flight Research Center INTRO

10、DUCTION The concept of manned entry vehicles capable of performing horizontal landings has been the subject of numerous theoretical and experimental studies. Among the many entry configurations studied, extensive wind-tunnel tests were performed to develop an entry shape designated the HL-10 (refs.

11、1to 5). In conjunction with these tests, a full-scale HL-10 lifting body vehicle was constructed for use in flight tests through the subsonic, transonic, and supersonic Mach number regions below 2.0. These flight tests are being performed to define the handling characteristics and the landing capabi

12、lity of the vehicle and to confirm the theoretical and wind-tunnel predictions of its stability, control, and performance characteristics. This paper defines the lift and drag characteristics of the HL-10 vehicle in four configurations over a Mach number range of 0.35 to 0.62 and at angles of attack

13、 from 5“ to 26“. The flight results, where applicable, are compared with full-scale and small-scale wind-tunnel results and the flight results obtained on an earlier manned lifting body entry vehicle, the M2-F2 (ref. 6). SYMBOLS Physical quantities in this report are given in the International Syste

14、m of Units (SI) and parenthetically in U. S. Customary Units. The measurements were taken in U. S. Customary Units. Details concerning the use of SI, together with physical constants and conversion factors, are given in reference 7. nondimensional cross-sectional area, perpendicular to the vehicle l

15、ongi tudinal axis a2 longitudinal acceleration, ratio of net aerodynamic force along the vehicle longitudinal axis to vehicle weight, g units an normal acceleration, ratio of net aerodynamic force normal to the vehicle longitudinal axis to vehicle weight, g units v Provided by IHSNot for ResaleNo re

16、production or networking permitted without license from IHS-,-,-II Ill1 11l111l11l1l1l1111111ll1l1m11lIl1111111111 II I1 b vehicle span, meters (feet) D CD drag coefficient, -CIS dcD 2 drag-due-to-lift factor dCL L cL lift coefficient, -CIS lift-curve slope per degree C variation of lift coefficient

17、 with elevator deflection, , per degree L6e cN Wannormal-force coefficient, -CIS cX Wa2axial-force coefficient, -CIS acL-D drag force along flight path, newtons (pounds) gravitational acceleration, 9.8 meters/second2 (32.2 feet/second2) L lift force normal to flight path, newtons (pounds) -L lift-dr

18、ag ratioD M free-stream Mach number M indicated Mach number AM Mach-number error, M - M NRe Reynolds number, based on vehicle length 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-P p AP cl P S W X-2 CY ACY4 ACYP ? 6e 6sb CJ corrected static press

19、ure, newtondmeter2 (pounds/foot 2) indicated static pressure from nose boom, newtons/meter2 (pounds/foo) position error in static pressure, p - 1, newtons/meter2 (pounds/foot2) dynamic pres sure, newtondmeter (pounds/foot2) 2reference area, body planform, meters2 (feet ) vehicle weight, kilograms (p

20、ounds) ratio of distance from nose of vehicle to an arbitrary point along longi tudinal axis to total vehicle length true angle of attack, am + ACY+ ACY + ACY + AaC, degreesP 4 E measured angle of attack, degrees angle of attack correction at 0“ angle of attack, due to angular difference between nos

21、e-boom incidence and the vehicles longitudinal axis, degrees angle-of-attack correction for effect of pitching rates on angle-of-attack vane , degrees angle-of-attack correction for nose-boom bending due to normal force, degrees angle-of-attack correction for effect of upwash factor on angle-of-atta

22、ck vane, (x)A% am, degrees elevon deflection, degrees speed-br ake deflection , degrees root-mean-square error 3 8 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Subscripts: max maximum mean average between right and left elevon deflections min mini

23、mum VEHICLE DESCRIPTION The HL-10 is a wingless. lifting configuration with a delta planform and negative camber . Heating was not a problem at the low Mach numbers of these flight tests. therefore aluminum was the primary material used to construct the vehicles semi monocoque structure. The pertine

24、nt physical characteristics of the vehicle are pre sented in table 1. and photographs are shown in figures l(a) and (b). TABLE 1. PHYSICAL CHARACTERISTICS OF THE HL-10 VEHICLE Body . Reference planform area. meters2 (feet2). Length. meters (feet) Span. meters (feet) biAspect ratio (basic vehicle). s

25、 . Weight. including pilot. kilograms (pounds) Center of gravity. percentage of reference length . Base area: Configuration A and B. meters2 (fee). Configuration C. meters2 (feet21 Configuration D. meters2 (fee) Elevons (two) -Area. each. meters2 (feet2) Span. each. parallel to hinge line. meters (f

26、eet) . Chord. perpendicular to hinge line: Root. meters (feet) . Tip. meters (feet) Elevon flaps (two) -Area. each. meters2 (feet2) Span. each. parallel to hinge line. meters (feet) . Chord. perpendicular to hinge line: Root. meters (feet) . Tip. meters (feet) Vertical stabilizer (one Area. meters2

27、(feed)- . Height. trailing edge. meters (feet) Chord: Root. meters (feet) . Tip. meters (feet) Leading-edge sweep. degrees . Rudders (two) -Area. each. meters2 (feet2) Height. each. meters (feet) Chord. meters (feet) Outboard tip-fin flaps (two) -Area. each. meters2 (feet2) Height. hinge line. meter

28、s (feet) Chord. perpendicular to hinge line. meters (feet). Inboard tip-fin flaps (two) -Area. each. meters2 (feet2) Height. hinge line. meters (feet) Chord. perpendicular to hinge line. meters (feet). 4 14.9 (160) 6.45 (21. 17) 4. 15 (13.6) 1. 156 2722 (6000) 51.8 1.38 (14. 83) 1.57 (16. 98) 2.71 (

29、29. 13) 1.00 (10.72) 1.09 (3.58) 0.59 (1.93) 1.24 (4. 06) 0.70 (7.50) 1.09 (3.58) 0.48 (1.58) 0.80 (2.63) 1.47 (15. 8) 1.53 (5.02) 1.32 (4.32) 0.60 (1.97) 25 0 . 41 (4.45) 1.26 (4. 12) 0 . 33 (1.08) 0.35 (3.77) 1.37 (4.50) 0.26 (0.84) 0.23 (2.48) 1.01 (3.31) 0.23 (0.75) Provided by IHSNot for Resale

30、No reproduction or networking permitted without license from IHS-,-,-(b) Gemdown. Fre 1. HL-IO vehicle. Figure 2 is a three-view drawing of the vehicle with dimensions and control sur faces specified. The elevons, which are the primary control surfaces, provide both pitch and roll control; the rudde

31、rs , located on the center vertical stabilizer, provide directional control and function as speed brakes with symmetrical outward deflection. In addition to the primary control surfaces, secondary surfaces are located on the tip fins and the upper surfaces of the elevons. These secondary surfaces ar

32、e used to change the vehicle configuration during flight and are adjustable by the pilot. 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-ers and Elevon flaps !ed brakes I I,_ I Elevons Longitudinal axis and horizontal reference Figure 2. Three-vie

33、w drawing of HL-I 0 vehicle. (Dimensions in meters (feet). Configurations A and B (fig. 3)were designed for maximum vehicle stability in the subsonic Mach number region (M 0.6). Configuration A was used during the full-scale wind-tunnel tests (ref. 8) and the initial flight test. -Speed brakes, 0“ R

34、udders and speed brakes I Tip fin Inboard -Outboard 21 Speed brakes, 0“ -Speed brakes, 0“ -Speed brakes, 8“ However, during the initial Configuration A -Outboard, 0“c3 Inboard, 0“ Elevon flap, 0“ Tip-fin flaps Configuration B Outboard, 0“c3 C Inboard, 0“ Elevon flap, 0“ Tip-fin flaps Configuration C

35、 Outboard, 4.5“c3 -Inboard, 5“ Elevon flap, 3“ Tip-fin flaps Confiiuratio Dc5 Inboard, 30.5“ Elevon flap, Up Tip-fin flaps Figure 3. HL-I 0 secondmy control suMaces in alternate configurations. 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-,:.- /

36、 Section drawing Cross-section view lnmrd ,-original contour Drooped Figure 4. HL-I 0 tip-fin modification. flight test, a severe flow disturbance was encountered on the vehicles upper surface. To alleviate the control and performance problems caused by this flow disturbance, the leading edges of th

37、e tip fins were drooped (fig. 4) to divert additional flow over the vehicles upper surface. This modification is the only physical difference between configurations A and B (fig. 3). After preliminary tests with configu ration B, an interim configuration (C) was used. For this configuration small ch

38、anges were made in the deflections of the second ary control surfaces (fig. 3) which increased the vehicles usable angle-of-attack range but did not significantly alter its longitudinal-stability char acteristics . Subsequent flight tests in the subsonic Mach number region (M 0.6) were made with con

39、figuration C. To alleviate stability problems encountered at the higher Mach numbers, the secondary control surfaces were deflected significantly (fig. 3, configuration D). Deflecting these surfaces increased the base area of the vehicle and thus resulted in greater longitudinal stability at Mach nu

40、mbers above 0. 6 (transonic Mach number region). Figure 5 shows the variation of the nondimensional cross-sectional area of the vehicle with percent of body length. The wing loading was approximately 183 kiloyams/ meter2 (37.5 pounds/foot2), based on the reference planform area of 14.9 meters (160 f

41、eet2). The center of gravity for these tests was approximately 51.8 percent of the reference length. /Configurations A and B 0 -X 1 Figure 5. HL-I 0 cross-sectionul-area distribution. 7 I . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TEST CONDI

42、TIONS Flight The flight lift and drag results presented were obtained during glide flights of the HL-10 vehicle with the landing gear up, following launch from a B-52 airplane. The data were obtained at altitudes below 13,700 meters (45,000 feet) and at Mach numbers between 0.35 and 0. 62. The vehic

43、le angle of attack was varied from 5“ to 26“, and the Reynolds numbers ranged from 25 X 106 to 62 X 106, based on the vehicle length of 6.45 meters (21, 17 feet). Wind Tunnel Full scale.- Prior to the flight tests of the HL-10 vehicle, wind-tunnel tests were conducted with the flight vehicle in the

44、NASA Ames Research Centers 40- by 80-foot wind tunnel (ref. 8). The data were obtained with the tip fins in the original contour (fig. 4), thus they are compared with the flight data obtained from the vehicle before the tip-fin leading edges were modified. Small scale.- A 0. 063-scale model of the H

45、L-10 vehicle was tested in the NASA Langley Research Centers high-speed 7- by 10-foot wind tunnel (ref. 9). Tests were made with the model in configuration A (unmodified tip-fin contours), Byand D (modified tip-fin contours). The wind-tunnel tests were conducted over a Mach number range of 0.35 to 0

46、.9. The test Reynolds number varied from 2.7 X 106 to 4. 0 X 106, based on the model length of 0.403 meter (1.322 feet). Base pressure measurements were obtained during the small-scale wind-tunnel tests. The effects of sting interference on the model base pressures and on the flow over the surfaces

47、ahead of the base were assumed to be negligible, although adequate wind-tunnel and flight base-pressure measurements have not been compared to sub stantiate this assumption. The wind-tunnel drag results were adjusted for approximately an order-of magnitude difference in Reynolds number between the f

48、light and model tests. This adjustment was derived from the Karmdn-Schoenherr flat-plate relationship modified for compressibility effects by the method of Sommer and Short (ref. 10). The resulting v increment of drag was applied to the model values of CD and -L as a constant re-D duction of drag co

49、efficient of 0. 0035 and are shown in the following discussion as an adjustment to the small-scale wind-tunnel results. Three-dimensional and lift effects on this increment were assumed to be negligible; similarly, the viscous effects attrib utable to scale differences on parameters other than drag were not consi