NASA NACA-RM-L53F12-1953 The calculated and experimental incremental loads and moments produced by split flaps of various spans and spanwise locations on a 45 degrees sweptback win.pdf

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1、SECURITY tNFORMATION :;G RESEARCH MEMORANDUM THE CALCULATED AND EXPERIMENTAL LTXREMENTAL LOAD3 AND MOMENTS PRODUCED BY SPLIT FLAPS OF VARIOUS SPANS AND SPANWLSE LOCATIONS ON A 45 SWEPTBACK WING OF ASPECT RATIO a By H. Neale Kelly * “r Langley Aeronautical La tF,T! ;:; ;?J;,$;. Langley Field, Va. NAT

2、IONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON September 4, 1953 ?“.i :. -. - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.B NACA RM L53Fl2 NATIONAL ADVISORY COEIMIm FOR AERONAUTICS RESEARCH bmlomuM THE CALCUU AND MpERrmw IMCREMENTAL LOADS A

3、ND MO“S PRODWED BY SPLIT FLAPS OF VARIOUS SPANS AND SPANWISE LOCATIONS ON A 45O ENEPTBACK WING OF ASPECT RATIO 8 By R. Neale Kelly SUMMARY The incremental lift and pitching moments produced by 20-percent- chord split flaps of various spans at various spanwise positions on two 45 sweptback wings of a

4、spect ratio 8.02 have been obtained by pressure- distribution tests in the Iangley 19-foot pressure tunnel. These data of 4,000,000 and a Mach number of 0.19. were obtained in the linear angle-of-attack range at a Reynolds number - The experimental data indicated that inboard flaps were far more eff

5、ective in prducing lift than outboard flaps (a 20-percent-span fnbmd flap produced approximately twice the increment in lift produced by a 40-percent-span outbmrd flap). Furthermore it was found that, in con- trast to the case for straight wings, the flap lift effectiveness (%) and the chordwise cen

6、ter of pressure of the incremental loads produced by full-span flaps on sweptback wings vary along the flap span. Comparison with the experimental data indicated that the procedure of NACA Technical Note 2278 can be used to predict the integrated incre- ments in lift and wing-root bending moment pro

7、duced by flaps on high- aspect-ratio, highly sweptback wings with fair accuracy. Probable causes of the deviations of the calculated loadings froan the experlmental have been discussed. For these wings the accuracy of the incremental pitching moment corn- puted by the method outlined in NACA Wartime

8、 Report L-164 is dependent primarily upon the accurate prediction of the spanwise load distribution. The spanwise variation of the chordwise center of pressure of the load produced by the longer span flaps could, by a simple modificatfon of the method, be closely approximated. Provided by IHSNot for

9、 ResaleNo reproduction or networking permitted without license from IHS-,-,-2 IN“I!RODUCTION NACA RM L53Fl2 A knowledge of the magnitude of the effects of flap geometry and position on the span loading and pitching-mcanent characteristics of a wing is required in the aerodynamic and structural desig

10、n of aircraft. Theoretical methods such as reference 1 are available for predictlng the loading produced by a deflected flap on straight and swept wings and the semiempirical method of reference 2 is available for approximating the incremental twisting and pitching mments. Because of the lack of lar

11、ge- scale experimental data, the applicability of the methods of references 1 and 2 to vugs with large amounts of sweep and relatively high aspect ratio, such as have been proposed for long-range, high-speed bombers, has not been ascertained. A general low-speed investigation is being made in the La

12、ngley 19-foot pressure tunnel on two 45O meptback wings having aspect ratios of 8.02 and taper ratios of 0.45. One of the wings is untwisted and incorporates an NACA 631A012 airfoil section in the free-stream direction; the other employs the same thickness distribution, but contains the calculated a

13、mount of twist and camber required to produce an elliptic span load dis- tribution and a uniform chordwise distrfbution at a lift coefficient of 0.7 and a Mach number of 0.9. As part of this investigation tests have been made, through the linear angle-of-attack range at a Reynolds number of 4,000,00

14、0 and a Mach number of 0.19, on the wings equipped with 20-percent-chord split flaps of various spans at various spanwise positions. Pressure data have been obtained Fn these .tests at seven spanrlse stations by means of orifices alined in the free-stream direction along the wing and flap surfaces.

15、The present paper contains the results of these tests and affords a ccanparison in a previously unchecked aspect-ratio-sweep range of the incremental lift and pitchug moment calculated by the methods of refer- ences 1 and 2 with experimental data. Results of other phases of the general investigation

16、 may be found in references 3 to 8. wing lift coefficient, C Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L53FI.2 0 3 . Cl section lift coefficient, b wing-root bending-moment coefficient, r s,* LOB elmo (Su - St ) d($) - cm section pitchi

17、ng-maPnent coefficient about c/4, c (su - ,)(0.25 - E) d(5) + cm wing pitching-mmnent coefficient about c /4, 1.0 C 2 s, CmC/4 x d(F) C sect ion p itching-manent coeff fc ient about c 14 , % 14 xc I /4 c, + - C cz I a wing lift-curve slope, - dc 2 da Provided by IHSNot for ResaleNo reproduction or n

18、etworking permitted without license from IHS-,-,-4 b C C C - S H P 9 a6 P V % X X CP XC /4 Y z a e wing span local chord parallel to plme of symmetry mean aerodynamic chord, ,so1*, c2d(F) mean geometric chord, .Ex pressure coefficient, - H-P b 9 free-stream total pressure local static pressure free-

19、stream amamic pressure, $ pV2 c flap lift effectiveness, -I- dcz dCZ d6 dar density of air free-stream velocity wing area NACA RM L53F12 longitudinal distance from local leading edge measured parallel to chord plane and plane of symmetry center of pressure of lding produced by flaps, fraction of loc

20、al chord longitudiaal distance from c /4 to c/4 lateral distance frm plane of symmetry measured perpendicular to plane of symmetry vertical distance frcm chord plane measured perpendicular to chord plane angle of attack of root chord gemetric angle of twist of any section referred to the root chord

21、(negative if washout) - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L53Fl2 5 . A increment produced by flaps . plane of symmetry 6 flap deflection angle measured in B plane parallel to the Subscripts: U upper surface 1 lower surface f for

22、ward of maximum thickness r rearward of msximMl thickness MODEL AND TESTS Experimental data presented in the present paper were obtained from tests of two wings of similar plan form. Each wing had an aspect ratio of 8.02, a taper ratio of 0.45, and 45 sweepback of the quarter-chord line. One wing wa

23、s untwisted and embodied HACA 631A012 airfoil sections in the free-stream direction; the other, which employed the amount of camber and twist determined by the method of reference 9 requbed to pro- tribution at a lift coefficient of 0.7 and a Mach number of 0.9, utilized the same thickness distribut

24、ion about a modified a = 1.0 mean line. The untwisted, symmetrical wing and the 80-percent-chord line (twist refer- ence axis) of the twisted and cambered wing had no dihedral. Additional geometric information can be obtained from figure 1 and references 3 and 4. - duce an elliptic span load distrib

25、ution and a uniform chordwise load dis- Both wings consisted of a solid steel core to which 8 bismuth-tin alloy was bonded. Surface orifices, distributed as illustrated .rr fig- ure 2 (wing orifices are listed in refs. 4 and 7) were provided for meas- uring the pressures on the left half-wing. Tubes

26、, leading from the ori- fices, embedded in the bismuth-tin alloy were brought out through the 20-percent-semispan station of the right half-wing. These tubes were conducted through a feiring (as sham in figs. 2 and 3) to multitube manmeters. Flaps used in the present investigation were of the split

27、type and were 20 percent of the local wing chord measured pezallel to the plane of symmetry. The flaps were constructed of sheet steel with provisions made for measuring pressures (see fig. 2 for locations) and were attached by means of steel angle block to the under surface of the wings. Flaps on t

28、he twisted and cambered wing were deflected 11.3 relative to the Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 NACA RM L53F12 local wing-chord plane while those on the untwisted, symmetrical wing were deflected 52.2O relative to the lower surface

29、 of the wing. Nl-span flaps, 60-percent-span inboard flaps, 40-percent-span outboard flaps, and 20-percent-span flaps at various spanwise locations were tested on the twisted and cambered wing. A 60-percent-span flap was tested on the untwisted, symmetrical wing. Pressure data were obtained through

30、an angle-of-attack range frm -2.60 to 4.8O for the twisted and cambered wag and *can 0.6 to g.Oo for the untwisted symmetrical wing. All tests reported herein were conducted on smwth Langley 19-foot presaure tunnel at a tunnel pressure of 2r atmospheres. The Reynolds nlonber, based on the mean and t

31、he corresponding Mach number of the tests were 4.0 respectively. 3 REDETION AND CaRRECTION OF DATA models in the approximately aerodynamic chord, x 106 and 0.19, Section lift and pitching-moment data were obtained by numerical integration of the chordwise pressure distributions. . All data in the pr

32、esent report have been corrected for Jet-boundary interference and airstream-angle variation in the region occupied by the models. More detailed discussions of theee corrections may be found in references 4 and 7. No corrections were applied to take into account the spanwise variation of the jet-bou

33、ndary-induced angle or the model twist due to air load. For all calculations on the twisted and cambered wing involving the flap deflection, a modified flap deflection was used. This modified flap deflection took into account the use of a flat flap on the highly cambered lower surface of the wing an

34、d was effectively the flap deflection relative to the larer surface of the wing. This modified flap deflectfon was deter- mined by the use of GLauerts thfn-airfoil theory and its variation along the span is presented Fn figure 2. RFVIEW OF ANALYTICAL METBODS Incremental Spanwise bading Any of the ex

35、isting methods for calculating the spanwise loading for swept wings could be utilized to calculate the incremental loading produced Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L53Fl.2 7 . by a partial-span flap provided sufficient spanwis

36、e control points were used in the solution to describe accurately the abrupt changes in the loading encountered at the end of the flap. The large number of control necessitate the expenditure of an excessive amount of cmputing time. An alternate procedure which makes possible the use of a small numb

37、er of con- trol points in the solution is presented in reference 1. - points required would, however, cmplicate the mathematical solution and In order to utilize a solution involving few control points in the calculation of the incremental loading produced by flaps, the discontinu- ous twist distrib

38、ution produced by such flaps must be approximated by a fictitious twist distribution which will yield the correct loading at these control points. Since this fictitious twist distribution Is assumed to be independent of plan-form effects the relatively simple, exact solu- tion for partial-span flaps

39、 on the zero-aspect-ratio wing (ref. 10) can be utilized in its calculation. Using the procedure outlined In refer- ence 1, for any given method for calculating the spanwise loading on straight and swept wings, a fictitious twist distribution can be developed that will permit the calculation of the

40、incremental loading produced by flaps without the use of an excessive number of control points in the solution. Reference 1 uses the tuist distribution determined for the zero- aspect-ratio wing in the procedure of reference 11 which is based on the simplified Weissinger 7-control-point solution of

41、Prandtls lifting sur- face theory to calculate the loading at four spanwise positions across the semispan (2y/b = 0, 0.383, 0.707, and 0.94). A special interpo- lation formula is also provided for the interpolation of the loading at four intermediate positions (2y/b = 0.195, 0.556, 0.831, and 0.g81)

42、. For fractional wing-chord flaps the full wing-chord flap loading obtained by this procedure is modified by the use of the two-dimensional flap lift effectiveness for the flap chord and deflection used. Incremental Pitching Moment According to reference 2, the center of pressure 0. the incremental

43、lift load produced by flaps over the flapped part of a wing x (relative to the leading edge) can be determined with satisfactory accu- cpf racy by the relation aO +Pfl = a In which a is the lift-curve slope and the subscript o pertains to the two-dimensional airfoil. The center of pressure over the

44、unflapped portion is then approxi- mated by a line faired between the 0.40-chord station at the flap end to the 0.25-chord station at a point 30 percent of the semispan from the flap u Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 rl NACA RM L53F

45、l.2 end; the center of pressure is thereafter assumed to fall along the 0.25-chard line 86 illustrated in the following sketch: Flap span 0.30 b/2 line to curve 0.40 The variation of the chordwise center of pressure thus determined is then used in conjunction with the moment arm produced by sweepbac

46、k and span- wise position hav- ever, at the outer spanwise stations, as the negative lift range is Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-!B NACA RM L53Fl2 9 approached, there is a strong positive trend of the local pitching moment for all c

47、onfigurations tested. Examination of the pressure distributions seems to indicate that this positive trend is due to flow separation over the lower surface induced by the large negative angles of attack at which the section is operating (see twist distribution, fig. 1). In contrast to the flap lift

48、effectiveness CY6 for strafght wings, which is for all practical purposes the same at all spanwise positions as that determined two dimensionally (as illustrated by the experimental data presented in ref. 12) the lift effectiveness for full-span flaps on the present sweptback wing varies considerabl

49、y along the flap span. In order to illustrate this difference more clearly, the spanwise variatfons of % for full-span flaps on a straight wing and the sweptback wing have been presented in figure 5. Spunrise Loading The calculated incremental span loadings produced by deflected flaps presented in figure 6