NASA NACA-RM-L51K26-1952 Force and pressure investigation at large scale of a 49 degrees sweptback semispan wing having NACA 65A006 sections and equipped with various slat arrangem.pdf

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1、RESEARCH MEMORANDUM FORCE AND PRESSURE INVESTIGATION AT LARGE SCALE OF A 4Q0 SWEPTBACK SEMISPAN WING HAVING NACA 65A006 SECTIONS AND EQUIPPED WITH VARIOUS SLAT ARRANGEMENTS By Stanley Lipson and U. Reed Barnett, Jr. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON January 29, 1952 Provided by

2、IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-* NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS RESEARCH MEMORANDUM FORCE AND PRESSUFE INVESTIGATION AT LARGE SCALE OF A 49O SWEPTBACK SEMISPAN WING HAVING NACA 65006 SECTIONS AND EQUIPPED WITH ,VARIOUS SLAT ARRANGEM

3、ENTS By Stanley Lipson and U. Reed Barnett, Jr. An investigation has been conducted to determine the effect of varying the span and deflection angle of a 15-percent-chord slat on the longitudinal aerodynamic characteristics of a semispan wing having 49.1 of 0.586, and incorporating NACA 65006 azrfoi

4、l sections streamwise. In addition to force measurements, chordwise pressure distributions were obtahed on the wingand extended slat with and without a deflected trailing-edge flap. The tests were conducted in the Langley f“l-scale tunnel with the greater part of the data being obtained at a Reynold

5、s -number of 6.1 x lo6 and at a Mach number of 0.10. . of sweepback at the leading edge, an aspect ratio of 3.78, a taper ratio I The results indicate that, from static longitudinal stability considerations, a slat span of 0.50 wing semispan was the most effective, for the subject wing, of the confi

6、gurations investigated; that slat spans shorter than 0.625 wing semispan had no effect on maximum lift; and, at a given lift coefficient, increasing the slat span and/or slat deflection up to 450 reduced the drag characteristics of the wing in the moderate- and high-lift range. INTRODUCTION The usef

7、ulness of slats in improving the low-speed characteristics of sweptback wings has been demonstrated in several investigations of specific high-speed plan forms. (See, for example, references 1 and 2. ) Inasmuch as the flow characteristics for a sweptback wing change -considerably with variations in

8、wing sweep and airfoil profile, the stall-control-device requirements (both aerodynamic and structural) for sweptback wings will also vary with wing geometry. c Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 In order to augment the limited amount

9、of available low-speed, large-scale, slat data on swept wing8, an investigauon was conducted in the Langley full-scale tunnel of a 15-percent-chord slat on a semi- span .lo sweptback wing. The wing has an aspect ratio of 3.78 and incorporates NACA 65006 airfoil sections streamwise. The longitudinal

10、force characteristics- of the wing were obtained for several slat spans and deflection angles. In addition, chordwise pressure distributions were determined from pressure orifices located on the wing and on the extended slat. Most of the data were measured at a Reynolds number of 6.1 x 106 and at a

11、Mach number of 0.10. . and the the CL COEFFICIENTS AND SYMBOLS The test data are presented as standard MACA coeffidents of forces - momenta. The data are referred to a set of axe8 coinciding with wind axes, and the origin was located at the .quarter-chord point of mean aerodynamic chord. - _ . “ max

12、imum ltft coefficient drag coefficient (Twice mod profile drag coefficient pitching-moment coefficient about the quarter-chord point of . .“ - the mean aerodynamic chord slat-section normal-force coefficient (.io pr d(%$ aspect ratio (b2/S) twice model span, feet local wing chord measured parallel t

13、o plane of symmetry, feet local slat .chord measured parallel to plane of symmetry, feet Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-* NACA RM 5x26 - 3 - - C local. Wing chord measured perpendicular to 0.50 line (see c fig. I), feet c f local tra

14、iling-edge flap chord measured perpendicular to 0.50 line, feet C s local. slat chord measured perpendicular to 0.50 line, feet - C mean aerodynamic chord, feet Pr lowei wing surface - upper wing surface P pressure coefficient e ipj P local static pressure, pounds per square foot - PO free-stream st

15、atic pressure, pounds per square foot - Q free-stream dynamic pressure, pounds per square foot R k$ Reynolds number S twice model area, square feet V free-stream velocity, feet per second X chordwise coordinate parallel to plane of symmetry, feet X chordwise coordinate measured perpendicular to 0.50

16、 line, feet Y spanwise coordinate perpendicular to plane of symmetry, feet YCP 3 spanwise location of the wing center of pressure, percent b a angle of attack, degrees 6s angle of deflection of slat, degrees P mass density of air, slugs per cubic foot F L. P coefficient of viscosity, slugs per foot

17、second Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 - NACA RM 5126 . . MODEL The wing was tested as a semispan configuration mounted on a reflection plane in the Langley full-scale tunnel as shown in figure 2. A complete description of the refle

18、ction-plane construction and the wind-tunnel flow characteristics in its vicinity are presented in reference 3. The wing had 49.1 of sweepback at the leading edge., an aspect ratio of 3.78, a taper ratio of 0.586, and no geometric dihedral or- twist. The streamwise airfoil section was an NACA 65006

19、section with the extreme tip of the wing-being half of a body of revolution based on the sane section ordinates. The wing plan form and some of the more pertinent dimensions are presented in figure 1. As may be noted from the wing layout of figure 1, a special significance is attached to the 0.50 li

20、ne. The mounting system of the subject wing was designed such that the sweepback angle of the wing may be varied and the pivot point of the arrangement is located on the 0.50 line. The chotce of this particular chord line for the pivot point was from mechanical rather than from aerodynamic considera

21、tions. The 0.9 line was then emgloyed 8s the reference chord line for layout purposes of the flap dimensions, preasure-tube installations, and SO forth, since peccentages of chord lengths normal to the 0.50 line remain constant regardlees of the angle of sweepback at which the wing- is being tested.

22、 Details of the arrangement employed for the investigation of a 15-percent-c slat are presented in figure 3. The ordinates-“of the slat are derived from those of the wing airfoil so that the slat could feasibly be retracted into a wing of the dimensians tested herein. For the present investigation,

23、howev.er, the slat was-mt constructed as an integral part of the wing and is mounted directly onto the unmodified basic-wing leading edge wi-th the slat brackets alined normal to the wings leading edge. The slat was composed of several individual spanwise segments so that slat spana of 0.250b/2, 0.3

24、75b/2, 0.500b/2, 0.625b/2, 0.750b/2, and 1.000b/2 could be obtained. All the slat tests were conducted with the outboard end of the slat located at the wing tip. The minimum chordwise clearance between the slat and wing, and the distance of the slat nose ahead of the wing were selected from the slat

25、- positioning results presented in reference 4 and were held constant when the slat angle , defined in figure 3, was varied. As shown in figures 1 and 3, the trailing-edge flap employed for the investigation had a 0.25 chord, a span of 0.469b/2, and a deflection angle of 45O measured normal to the 0

26、.50 line. “ . “ . . “- :. “. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM 5x26 - 5 . 1 Flush surface ptatic-pressure orifices were installed in the slat in three chordwise rows in the stream direction and in the wing in two chordwise rows

27、in the stream direction and seven chordwise rows normal to the 0.50 line. The general spanwise arrangement of these pressure orifices is shown in figure 4 and their chordwise locations are presented in table I. TESTS AND CORRECTIONS Both force and pressure measurements were obtained during the inves

28、tigation. The force-data phase of the program was conducted over an angle-of-attack range from approximately -2O to 31. The slat param- eters varied during the test program were the slat deflection angle (with the slat span held constant at 1.00b/2) and the slat span (for a slat deflection angle of

29、45O). For the two test configurations, the basic wing and the wing with the 0.50b/2 span slat deflected 45O, the test Reynolds number was varied- from 2.9 x lo6 to 6.1 x lo6. The remainder of the test program was conducted only at a Reynolds number of 6.1 X 10 and a Mach number of 0.10. + 6 3. Chord

30、wiae pressure distributions and tuft surveys were obtained at four representative angles of attack for the following conditions: . (a) basic wing, (b) wing with 0.50b/2 span slat deflected 450, and (c) configuration (b) with the trailing-edge flap deflected. The jet-boundary corrections applied to t

31、he force data, as discussed in reference 3, presented herein were based on the method presented in reference 5 and were added to the uncorrected results. La = -0.84C ED = -0 .01281cL2 Ern = -0.00427C The data have also been corrected for the effects of blockage and stream angle. RESULTS A.ND DISCUSS

32、ION I The results of the force tests are presented in figures 5 to 8, - andloads on the slat in figures 14 to 16. the wing flow surveys in figure 9, the pressure .Ustributions over the wing in figures 10 through 13, and the chordwise pressure distributims Provided by IHSNot for ResaleNo reproduction

33、 or networking permitted without license from IHS-,-,-6 NACA RM 5x26 Aerodynamic Characteristics Basic wing.- The effects of Reynolds number on the aerodynamic characteristics of the basic wing me presented in figure 5. The data exhibit the trends characteristic of the results obtained for a thin sw

34、eptback wing having the leading-edge separation vortex type of flow. The lift coefficient at which the increase in wing lift-curve slope (which is inherent with the separation-vortex type of flow) idtially occurs increases in magnitude with Reynolds number (fig. 5(a). The lift-curve slope measured t

35、hrough zero lift agrees w.ith the value predicted by using the Weissinger method (reference 6) The Ch “ obtained at R = 6.1 x lo6 was approximately I. 00, or about 0.02 higher . tlian the Cbx reached for a test Reynolds number of 2.9 x 10 6 . “ The increased tendency for the experimental drag curve

36、to deparrt “ “. from the theoretical curve CL2 in the low-lift range with a decrease in Reynold8 number shows the unfavorable influence that decreasing R exerts on the flow characteristics of the wing (fig. 5(b). . . A comparison of the drag data, on the basis of constant lift coefficfmrt, . indicat

37、es that CD is decreased by an increase in the test Reynolds number throughout the lift range above about CL = 0.3. 4 - “ . . 1 Varying the Reynolds number over the range.i,nvestigated does not appear to alter significantly the general shape of the pitching-moment and spanwise-center-of-pressure cweB

38、 (figs. 5(c) and ?(a). Effect of slat deflection angle.- The effect of the slat deflection “ angle was determined only for the full-span slat arrangement (fig. 6). An increase in Ch i8“obtained by increasing 8s up to the highest slat deflection angle tested (fig, 6(a) ). At 8s = 450, Chx is approxim

39、ately 1.17 as compared to about 1.00 for the basic wing. The improvement in the drag characteristics of the wing obtained “ by increasing the slat deflection may be illustrated by the fact that, - at a lift coefficient of 1.00, the drag of the wing with the full-span slat deflected 20 is approximate

40、ly 80 percerrt greater than that measured for the SS = 45O configuration. “. - Deflection of the full-span slat results in instability of the wing over the entire lift range from zero lift to maxfmm lift (fig. 6(c). Up to a moderately high lift coefficient the wing with the slat deflected is moderat

41、ely unstable and then, very abruptly, becomes highly unstable. The abrupt change in stability, similar to the characteritics obtained IX Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM 5x26 - 7 with the basic wing, is.caused by a loss of lift

42、 outboard, as indicated by the inboard movement of the spanwise center of pressure in figure 6(d) and by the decrease in the wing lift-curve slope in the same lift- coefficient range where this stability change occurs (fig. 6(a) ). Increasing 6s increases the magnitude of the CL at which the abrupt

43、change in stability occurs (fig. 6(c). For the wing with the full- span slat deflected 45O, the change in stability from the moderate to the hfgh lift range is approximately equivalent to a 0.G shift forward of the wing center of pressure. As in the case of the basic wing, a stable change in the pit

44、ching-moment characteristics is obtained at the stall for all of the deflected full-span slat configurations tested. Effect of slat span.- Figure 7 illustrates the influence of the span of the extended slat, 6s = 45O, on the aerodynamic characteristics of the wing. Increasing the slat span has no ef

45、fect on Cha for spans shorter than 0.625b/2 but produces significant increases in Cba for spans between 0.75b/2 and 1.00b/2. c As shown in figure 7(b), correspondingly greater decreases in drag I than 0.25b/2 are employed. are obtained in the moderate and high-lift range as compared to the character

46、istics of the basic wing as successively longer slat spans The general trend of the longitudinal stability and spanwise center of pressure.with lift coefficient (figs. 7(c) and 7(d), respectively) indicates a stability change for slat spans greater than 0.50b/2. For slat spans of 0.75b/2 and 1.00b/2

47、 (fig. 7( d) ) , the location of the span- wise center of pressure remains fairly constant over a large range of lift coefficients. Between CL of 0.2 and 0.5, however, the length of slat span does not appear to alter the spanwise location of $he center of pressure (fig. 7(d) and the difference in st

48、ability in.the 19w-lift range, shown in figure 7(c), is probably caused by a more forward chord- wise shift of the center of pressure due to the use of the longer slat spans. For both the basic wing and the configurations of slat spans greater than 0.50b/2, the wing becomes highly stable at Cba, whe

49、reas for the shorter slat spans, 0.2%/2 and 0.375b/2, this stable effect occurs initially at a CL about 0.1 lower than Ch. This effect may be correlated with the variations in spanwise location Of the Wing center of pressure near (fig. 7( d) ) . O.5Ob/2 slat span. - On the basis of the stability characteristics I shown in figur