NASA-TP-1889-1981 Wing-alone aerodynamic characteristics for high angles of attack of supersonic speeds《在超音速高攻角下的单机翼空气动力特性》.pdf

《NASA-TP-1889-1981 Wing-alone aerodynamic characteristics for high angles of attack of supersonic speeds《在超音速高攻角下的单机翼空气动力特性》.pdf》由会员分享，可在线阅读，更多相关《NASA-TP-1889-1981 Wing-alone aerodynamic characteristics for high angles of attack of supersonic speeds《在超音速高攻角下的单机翼空气动力特性》.pdf（191页珍藏版）》请在麦多课文档分享上搜索。

1、NASA Technical Paper 1889 - Wing-Alone Aerodynamic , Characteristics for High Angles of Attack I at Supersonic. Speeds Robert L; Stallings, Jr., and Milton Lamb - JULY 1981 . r / - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM

2、 NASA Technical Paper 1889 Wing-Alone Aerodynamic Characteristics for High Angles of Attack at Supersonic Speeds Robert L. Srallings, Jr., and Milton Lamb Langley Research Center Hamnpton, Virgitpia National Aeronautics and Space Administration Scientific and Technical Information Branch 1981 Provid

3、ed by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-SUMMARY An experiment has been conducted to determine wing-alone supersonic aero- dynamic characteristics at high angles of attack. The wings tested varied in aspect ratio from 0.5 to 4.0 and taper ratio from

4、 0 to 1.0. The wings were tested at angles of attack from -5O to 60 and Mach numbers from 1.60 to 4.60. Aerodynamic forces and moments and center-of-pressure locations were obtained by integrating pressure measurements over the wing surface. The longitudinal and lateral center-of-pressure locations

5、approached the wing-area centroids at the maximum test angles of attack. For angles of attack greater than approximately 30, the center-of-pressure locations were not sig- nificantly affected by Mach number. Increasing the aspect ratio resulted in a general increase in normal-force coefficient CN. I

6、ncreasing the taper ratio X from 0 to 0.5 resulted in an increase in CN but further increases in X had little effect on CN. Peak pitching-moment coefficients were measured at the approximate angle of attack at which free-stream pitot pressure was first measured on the windward surface in the wing-ap

7、ex region. At the maximum test angle of attack (60), pitching-moment coefficient was not significantly affected by either aspect ratio or taper ratio. INTRODUCTION The high-maneuverability requirements of missiles often necessitates flight at high angles of attack. At these high angles of attack, po

8、tential-flow methods and linear theories have very limited applications, and the missile designer generally resorts to semiempirical methods based on wind-tunnel data for prelimi- nary design purposes. However, there is a lack of a systematic data base of wing-alone forces and moments as a function

9、of aspect ratio, taper ratio, and Mach number at high angles of attack (ref. 1). This lack of a data base results, in part, from the difficulty associated with obtaining data unaffected by support interference at the higher angles of attack. In order to provide some of the needed data for high angle

10、 of attack, an experimental program was conducted at the NASA Langley Research Center using pressure models and a sting support system that was designed to minimize sting interference effects for the test range of angles of attack. Aerodynamic forces and moments and center-of-pressure loca- tions we

11、re obtained by integrating the pressure measurements. The wings tested varied in aspect ratio from 0.5 to 4.0 and in taper ratio from 0 to 1 .O. Angle of attack varied from -5O to 60 at Mach numbers from 1.60 to 4.60. Both aerodynamic forces, aerodynamic moments, and pressure data are presented and

12、discussed. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-SYMBOLS aspect ratio b wing span bps semispan of wing center-planar section (fig. 8) span of wing tip (fig. 8) btip CA axial-force coefficient (see appendix) cb bending-moment coefficient (se

13、e appendix) Cm pitching-moment coefficient (see appendix) normal-force coefficient (see appendix) Cn yawing-moment coefficient (see appendix) P - P, pressure coefficient, q cP - C mean aerodynamic chord FA L axial force centerline length (2 in computer-generated tables, tables to x) length of leadin

14、g edge (fig. 8) LLE LPS LTE M length of wing center-planar section (fig. 8) length of wing trailing edge (fig. 8) free-stream Mach number Mb bending moment %om MY MZ N nominal free-stream Mach number pitching moment yawing moment normal force static pressure P free-stream static pressure 2 Provided

15、by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1 11; 1 Pt ii li. free-stream stagnation pressure xc,w xc , AS xCP X XPS Y YPS Ytip z zc , AS free-stream dynamic pressure free-stream Reynolds number, per meter wing planform area free-stream stagnation tempera

16、ture wing thickness free-stream velocity longitudinal distance measured downstream from wing apex (fig. 5) longitudinal distance measured upstream from wing trailing edge, x=L-x value of x at wing-area centroid value of x at area centroid of element of planform area AS value of x at wing longitudina

17、l center-of-pressure location (see appendix) downstream distance from wing leading edge (fig. 8) downstream distance from forward edge of wing center-planar section (fig. 8) downstream distance from aft edge of wing center-planar section (fig. 8) perpendicular distance from wing centerline measured

18、in plane of wing (fig. 5) value of y at wing half-panel area centroid value of y at area centroid of element of area AS value of y at wing lateral center-of-pressure location for wing half panel (see appendix) spanwise distance from axis of symmetry on wing center-planar section (fig. 8) spanwise di

19、stance from outer edge of wing center-planar section (fig. 8) perpendicular distance from wing horizontal plane of symmetry (fig. 5) value of z at area centroid of element of area AS 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-+P value of z at

20、wing vertical center-of-pressure location (see appendix) a angle of attack AS element of planform area A sweep angle x taper ratio Subscripts: LE leading edge TE trailing edge APPARATUS AND METHODS Wind Tunnel and Test Conditions The tests were conducted in both the low and high Mach number test sec

21、tions of the Langley Unitary Plan Wind Tunnel, which is a variable-pressure continuous- flow facility (ref. 2). Asymmetric sliding-block nozzles lead to the test sec- tions and permit a continuous variation in Mach number from about 1.50 to 2.90 in the low Mach number test section and from about 2.3

22、0 to 4.70 in the high Mach number test section. The tests were conducted at angles of attack ranging from -5O to 60 for test conditions listed in the following table: , M Tt , Pt kPa K 1.60 352 249.26 4.60 1 137.99 3.50 98.44 2.86 68.47 2.16 339 54.63 6.56 x lo6 1 Since friction drag cannot be deter

23、mined from pressure measurements, and since the size of artificial roughness required to trip the boundary layer at the highest supersonic Mach numbers can distort the inviscid flow field, no attempt was made to artificially trip the boundary layer on any of the wings. 4 Provided by IHSNot for Resal

24、eNo reproduction or networking permitted without license from IHS-,-,-Models and Instrumentation The wings tested consisted of 10 models that had aspect ratios ranging from 0.5 to 4.0 and taper ratios ranging from 0 to 1 .O. Figure 1 (a) is a photograph of the models, figure l(b) shows a typical ass

25、embly consisting of one of the models and the dogleg sting, and figure 1 (c) gives the basic sting dimensions. The sting was designed to minimize support interference effects on the wing surface opposite the sting-attachment point. This opposite surface was instru- mented with pressure orifices. Loc

26、ations of the pressure orifices. for the 10 models are shown in figure 2. Since the wings are symmetrical about. the lon- gitudinal centerline, most of the pressure instrumentation was located on only half a wing panel. Two orifices were located on the opposite panel to insure flow symmetry. Also gi

27、ven in figure 2 are the basic model dimensions and the wing-area centroid locations. All wings had planform areas of 232.26 cm2, a maximum thickness of 1.27 cm, and had leading edges, tips, and trailing edges consisting of sharp wedges with a total angle of 30 measured in a plane perpen- dicular to

28、the edges. Since only one side of the wings was instrumented with pressure or if ices, windward and leeward measurements at a given angle of attack were obtained by testing the model assembly in the attitudes shown in figure 3. Shown in fig- ure 3(a) is the attitude of the assembly with the instrume

29、ntation windward and in figure 3(b) is the attitude of the assembly with the instrumentation leeward. The attitude of the assembly shown in figure 3(b) was obtained by simply rotat- ing the complete assembly 180 from the attitude shown in figure 3 (a) . Data Reduction The pressure measurements for b

30、oth the windward and leeward surfaces were reduced to coefficient form and are presented in tables I to X. Forces and moments were obtained by integrating the pressures over the windward and leeward surfaces using the equations shown in the appendix. These integrated values were combined to obtain t

31、he total forces and moments as illustrated in figure 4 for the case of normal-force coefficient. Since the angles of attack for the windward and leeward tests differed slightly (by 0.02O or less), the integrated forces and moments were evaluated at the nearest integer value of angle of attack by usi

32、ng a second-order Lagrangian interpolation before combining the results from the two surfaces. The angles of attack shown for the tabulated pressure measurements are the averages of the windward and leeward values. The moment-center locations and sign conventions are shown in figure 5. It should be

33、noted that since the forces and moments were evaluated for only half a wing panel, the reference areas used to determine the force and moment coeffi- cients were half the wing planform area, or 11 6.1 3 cm2. However, the values of CN, CA, Cm, xcp, and zcp would also be applicable for a complete wing

34、 panel as presented, since the flow is symmetrical about the longitudinal axis of symmetry . 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-RESULTS AND DISCUSSION A complete tabulation of the pressure measurements that were integrated over the ins

35、trumented surface to obtain the aerodynamic characteristics are presented in tables I to X. Presented in figure 6 is a comparison of CN, xcp, and ycp from the pres- ent tests with some previously published wing-alone data. The Hill data (ref. 3) are force-balance data from a sting-supported model. T

36、he Falunin et al. data (ref. 4) are from both force-balance and pressure measurements, and the Baker data (ref. 5) are from force-balance measurements of semispan wings attached to a reflection plane. The Nielsen et al. results (ref. 6) are a data base that was compiled from existing wind-tunnel dat

37、a including references 3 to 5. It should be noted that most of the referenced data were not directly available for the aspect ratio, taper ratio, or nominal free-stream Mach number indicated in figure 6 but were determined for this comparison by a linear interpolation of the available data with resp

38、ect to these variables. The comparisons shown in figure 6(a) are for a nominal free-stream Mach num- ber of 2.00, tR = 0.5, and X = 0. In general, good agreement is shown between the various sets of data with the exception of the reference 5 reflection-plane data at the larger angles.of attack. The

39、reference 5 data is believed to be influenced at the larger angles of attack by flow separation on the reflection plane that was induced by the large pressures on the wing windward surface. This effect (as shown in fig. 6(a) results in a reduction in normal force and an outboard movement of the late

40、ral center of pressure. The longitudinal center-of- pressure location is not significantly affected by this separation phenomenon. Agreement similar to that shown in figure 6(a) is also shown in figure 6(b) for Mnom = 3.00. The onset of the effect of separation for the reference 5 reflection-plane d

41、ata occurs at a lower angle of attack at Mnom = 3.00. Shown in figure 6 (c) are data comparisons for wings having A? = 1 .O and X = 0 at Mnom = 2.00. The present data are in good agreement with the refer- ence 4 data throughout the angle-of-attack range and with the reference 5 data for the angles o

42、f attack below the onset of the reflection-plane separation. The values of CN from both references 3 and 6 are generally greater than the other data presented. The extracted data base of reference 6 is based on the reference 3 data for the case of X = 0, with the other sources of data being used to

43、determine the effects of taper ratio. Therefore, the reference 6 data base would be expected to agree with the reference 3 data. The good agreement between the other three sets of data suggests, however, that the reference 3 data are suspect. Similar results are shown at Mnom = 3.00 in figure 6(d).

44、Data for wings with A? = 2.0 and X = 0 were only available from refer- ences 4 and 5 and these results are compared with the present data at nominal free-stream Mach numbers of 2.0 and 3.0 in figures 6 (e) and 6(f). With the exception of the reflection-plane data affected by separation, the data are

45、 generally in good agreement. Shown in figure 6 (9) are data comparisons for wings with R = 1 .O and X = 0.5 at a nominal free-stream Mach number of 3.00. For this case, data 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-/ for comparison purposes

46、 were available only from references 5 and 6, although, i as previously mentioned, the reference 6 data are an extension of the reference 3 data for X = 0. The trends shown in figure 6 (9) are basically the same as shown in the previous figures for X = 0. Forces and moments for the rectangular wings

47、 (x = 1 .O) were estimated from pressure distributions determined from oblique-shock and Prandtl-Meyer-expansion relations (shock-expansion method) and are compared with the measured data in figure 7. The estimations are shown for angles of attack up to 25O, which is the approximate angle of attack

48、for shock detachment on the leading edge at Mach 4.60. The results show generally good agreement between measured and cal- culated pitching-moment coefficients and axial-force coefficients; however, the calculated normal-force coefficients and bending-moment coefficients generally averpredict the me

49、asured values. The extent of this disagreement decreases with increasing aspect ratio since the two-dimensional assumption of the theory is more nearly satisfied at the higher aspect ratios. It should be noted that the measured and calculated axial-force coefficients were determined from pressure distributions and, therefore, do not include any f