NASA NACA-TN-4366-1958 The effects of an inverse-taper leading-edge flap on the aerodynamic characteristics in pitch of a wing-body combination having an aspect ratio of 3 to 45 de赫.pdf

1、-40NATIONALADVISORYCOMMITTEEFOR AERONAUTICSTECHNICAL NOTE 4366THE EFFECTS OF AN INVERSE-TAPER LEADING-EDGE FLAPON THE AERODYNAMIC CHARACTETICS IN PITCH OFA WING-BODY COMBINATION HAVING AN ASPECTRATIO OF 3 AND 45 OF SWEEPBACK ATMACH NUMBERS TO 0.92By Fred A. Demele and K. Harmon PowellAmes Aeronautic

2、al LaboratoryMoffett Field, Calif.WashingtonAugust 1958Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECHLIBRARYKAFB,NMTEE EWFECTSOF AN INVERSE-TAPER IJ3ADING-EIX3EFIAPON THE AERODYNAMIC CHARACTERISTICS IN PITCH OFA WING-BODY COMBINATION HAVTNG AN

3、ASPECTRATIO OF 3 and 45 OF SWEEPBACK ATMACH NUMBERS TO 0.92By Fred A. Demele and K. Harmon PowellSUMMARYAn investigationhas been made to detemnine the effects of aninverse-taperleading-edge flap on the drag ad on the static-longitudinalcharacteristicsof a swept-wing-bodyconibination. The wing had 40

4、 ofleading-edge sweepback, an aspect ratio of 3, a taper ratio of 0.4, andno camber or twist. However, with the flap deflected, the wing had acamber and twist distribution similar to that resulting from the incor-poration of conical camber in the forward portion of a plane wing. Thetests were conduc

5、ted over a range of Mach numbers from 0.25 to O.$E?at aReynolds number of 3.2 million, and over a Reynolds number range of3.2million to 15 million at a Mach number of 0.25 with flap deflections to 160.In the range of Mach numbers from 0.60 to 0.92, deflection of theflap resulted in significant dxag

6、reductions at lift coefficients of 0.2and greater. For optimum flap deflection, the maximum lift-drag ratioswere near the estimated maxinmms based on the assumptions of ellipticspan loading and full leading-edge suction. Slightly higher increases inmaximum lift-drag ratio were associated with optimu

7、m flap deflection thanwith conical csmber. At a Mach number of 0.25 and at a Reynolds nuniberof 15 million the flap was effective in reducing drag only at lift coeffi-cients above 0.55. In general.,the flap had little effect on the liftand static stability of the model.JNIRODUCTIONFor certain missio

8、ns of airplanes capable of supersonic flight, itmay be most economical to cruise at high subsonic speeds. Thus, it isimportast that the subsonic lift-drag ratios be maximized with minimumpenalty to the supersonic capabilities of the sirplane. Supersonic flightnecessitates the use of thin wings which

9、 are not conducive to highProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACA 4366aerodynamic efficiency at subsonic speeds. The usual leading-edge shapeof such wings causes separation to occur at a low lift coefficientand Gconsequentlyprevents th

10、e attainment, above that lift coefficient,of aneffective leading-edge suction force necessary for low drag due to lift.It was shown in references 1 and 2 that it is possible to attain very lowvalues of drag due to lift at subsonic speeds by incorporating conicalcamber over the forward portion of thi

11、n wings, even though such camberis designed for low supersonic speeds. Caniberingin this manner causesthe leading-edge suction pressures (i.e.,pressures lessstatic) to be distributed over a larger ar-a. Hence, toleading-edge suction effect required for low drag due tosures need not be as low as if t

12、hey were concentratedatand are therefore physically realizable. Although largemaximum lift-drag ratio at high subsonic speeds resultedthan free-streamproduce thelift, these pres-the leading edgeimprovements infrom camberinga wing in this mere small mi this gap was filled to pro-vide a smooth uper su

13、rface. The fuselage had a Sears-Haack shape offineness ratio 12.5. Geometry of the model and the equation of thefuselage shape are given in figure 2. .TESTS wLongitudinal force and moment data were obtained for flap deflectionsof 0, 4, 8.5, 12, and 160 throughout an angle-of-attackrange from -2to 20

14、, except at high llachnumbers where the angle limit was reducedbecause of tunnel power limitations. The major portion of the investiga-tion was made over a Mach number range from 0.25 to 0.92 at a Reynoldsnurfiberof 3.2X106, and over a Reynolds nudber range from 3.2XI.06to15x10e at a Mch number of 0

15、.25. In generslj the tests at a series ofMach numbers and constant Reynolds number were made with a 0.005-inchwire trip affixed to the upper and lower srfaces of the wing l/16-inchbehind the flap hinge line. The wire was removed-for tests at a seriesof Reynolds numbers and constant Mach number.The w

16、ire was employed to fix transition on the wing in an effortto maintain a skin friction of nearly constant maitude throughout theangle-of-attack range. The size of the wire was selected on the basisof the empirical results reported in reference 3. To verify that transi-tion was induced by the wire, u

17、se was made of a sublimationtechnique(see ref. 4) employing either acenaphthene or biphenyl in solutionwithpetroleum ether.*Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.uNACA TN 4366Static pessures were measured at thethe model to determine the t

18、est conditions5tunnel WSU in the region offor which the data may havebeen affected by 10CSL choking of the air stream at high Mach numbers.CORRECTIONSThe data have been corrected for tunnel-wall interference associatedwith lift on the wing, for blockage due to the yresence of the tunnelwalls, for ef

19、fects due to a streamwise static-pressure gradient, smd forlongitudinal force tares of the turntable on which the model was mounted.The method of reference 5 was used to evaluate theeffects. The resulting correctionswhich were added tothedueThearecoefficients are as follows:Act= 0.607 A% = 0.0083 %2

20、ACm= o.= cLCorrections to the data to account for the effectsWaLl interferencethe eagles mdof constrictimto the tunnel w: 1.005.80 1.010.85 .841 1.013.90 .884 1.019.92 a71 900 L 023A correctionwas applied to the drag to account for the drag forceon the model resulting from the tunnel streauwise stat

21、ic-ressuregradient.The vue of this drag coefficient correctionwas never greater than 0.0006.The corrections associated with dxag tare force due to aerodynamicforces on the eqosed surface of the turntable are given in the followingtable. No attmpt has been made to evaluate possible drag forces due to

22、interferencebetween the model and turntable.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 NACATN 4366M % tare0.25 0.0028.0028:2 .003285 .0033.0036:$ .0038RESULTS AND DISCUSSIONThe basic longitudinal characteristicsof the model are presentedgraphi

23、cally in figures 3 through 8 for severalReynolds numbers at a Machnumiberof 0.25, and in figures 9 through 14 for several Mach numbers at aReynolds number of 3.2x106. Selected drag and lift-drag characteristicsare presented as functions of Reynolds number in figures 15 and 16 and asfunctions of Mach

24、 number in figures 17 through 20. Figure 21 shaws theeffect of these parameters on the slopes of the lift and pitching-momentcurves. An index to these figures is presented as table II.Since the Reynolds numbers available at high subsonic speeds forthis investigationwere low comparedwith fulJ-scaLe v

25、alues, sm attemptwas made to fix the magnitude of the skin friction throughout the angle-of-a.ttackrange by fixing transitionnear the wing leading edge. Thus,the preponderance of data for the evaluation of the effects of Mach nun-ber (figs.9 through 11) was obtained for the wing with a wire tripaffi

26、xed to the upper and lower surfaces near the leading edge. Testswere also made with the wire off for flap deflections of 0 and 4(figs.12 through 14) in order to evaluate the effects on the aerodynamiccharacteristicsof using the wire to fix transition. A sublimationtechniquewas used in flow studies a

27、t high subsonic speeds and showedthat with the flap unreflected transition occurred close to the wire,whereas free transition occurred to the rear of the midchord line. How-ever, for most of the low-speed tests the transitionwire was not used(see figs. 6 through 8) since sublimationf16w studies indi

28、cated thatthe Reynolds numbers availablewere of sufficientmagnitude to causetransition to occur near the leading edge. Tests were also made withthe wire on, flap unreflected (figs.3 through 5), for the purpose ofevaluating the effects of Reynolds number on pressure drag for the con-dition wherein mo

29、st of the wing was inmersed in a turbulentboundarylayer. On the basis of sublimationflow studiesmade at low speeds andlow Reynolds numbers, free transition,while not clearly defined, appearedto occur well forward of the midchord position except for a region nearthe tip; the addition of the wire caus

30、ed transition to reeveclose to thewire.3.mi-.-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-IOWA TN 4366 7At Mach numbers of 0.90 and O.* partial choking of the wind tunneloccurred in the region of the model at the higher angles of attack.Dashed li

31、nes were used in fairing curves through basic data points forwhich a state of partisl choking of the wind tunnel was indicated.Drag CharacteristicsEffects of Reynolds number.- At low speeds and smll flap angles asignificant reduction in drag occurred at lift coefficients above 0.28 to 15x106 (figs.k

32、(a) andwith increasingReynolds number from 3.2xlO15). This phenomenonwas evidenced at zero flap deflectionwith andwithout the transitionwire Up to the lift coefficient at which maxhumlift-drag ratio occurred, the reduction in drag coefficientwas essentiallyconstant and is attributed to a reduction i

33、n skin friction with increasingReynolds number. This is demonstratedby the data of figure k(a) smd alsoby figure 16which shows, for zero flap deflection, the near attainment ofthe estimated maxixmm lift-drag ratio throughout the Reynolds number rangeof the test. Experimental.values of minimum drag c

34、oefficient for the planewing-body conibinationwere used for the estimated values of maximum lift-drag ratio, and it was assmd that the span loading was elliptical andthe leading-edge suction force was maximum. At higher lift coefficientsthe drag reductions accompanying an increase in Reynolds number

35、 (seefig. 15) are attributed to a greater effective leading-edge suction force.Apparently the low pressures required for attainment of fullleffectiveleading-edge suction force are not realized at low Reynolds numbers withthe flap at zero deflection. However, as indicated in figure 15, thereis (at lo

36、w Reynolds number) a progressive decrease in drag coefficient atconstant lift coefficient with increasing flap deflection, and the effectsof Reynolds number practically disappear. It is surmised that the csziberintroducedby deflection of the fla redistributes the suction pressuresover a larger regio

37、n. Thus the magnitude of the pressures for fullleading-edge suction (i.e.,pressures required for minimum drag due tolift) would be lower than if they were concentrated at the leading edge,and therefore these pressures are more nearly physically attainable.At Mach numbers of 0.60 and O. whereas, coni

38、-cal csmberwas more advantageous at speeds near a Mach number of 0.90.Lift and Pitching-Moment CharacteristicsExamination of the low-speed lift and pitching-moment data in fig-ure 6 and of the high-speed data in figure 9 revesls that, in genersl,deflection of the flap resulted in more nearly linear

39、pitching-momentcurves. It may be further noted that only a slight negative moment shiftG occurredwith flap deflection; consequently, the drag associated withtrinming these madnts would be mall. The data in figure 21 show thatthroughout the Reynolds nu?iberand Mach number range of the investigation*t

40、he slopes of the lift and moment curves near zero lift generally increasedslightlywith increasing flap deflection.CONCLUSIONSAn experimental investigation has been conducted to determine theeffectiveness of sm inverse-taperleading-edge flap in improvingprimarily the drag characteristics at high subs

41、onic speeds of awing-body combination having an aspect ratio of 3 smd 45 of leading-edgesweepback. The results can be summarized as follows:1. In the range of Mach numbers from 0.60 to 0.92, deflection ofthe flap resulted in significant reductions in drag at lift coefficientsof 0.2 ad greater. The m

42、sxlmum lift-drag ratios were nearly 20 percenthigher than those of the ane wing and were near the estimated maximums,based on the assumptions of elliptic span loading d full lesiking-edgesuction.2. Compared to conical csmber, the lesiling-edgeflap promotedslit3y larger gains in maximum lift-drag rat

43、io in the Mach number. rmge from 0.60 to 0.92.a71wProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-10 NACA TN43663. At low speeds and at a Reynolds nu?iberof 15 million the flap*was effective in reducing”drag coefficient only at lift coefficientsabove

44、 0.35. b4. In general, deflecting the flap resulted in little change inlift static stability.Ames Aeronautical LaboratoNationsl Advisory Committee for AeronauticsMoffett1. Hall, CharlesRatio Wings1953.Fieidj Cslif., MY 13, 1958REFERENCESF.: Lift, Drag, and Pitching Moment of Low-Aspect.at Subsonic a

45、nd supersonic Speeds. NACA RMA53A302. Ssmmonds, Robert I., ad Reynolds, Robert M.: The Effect of Conical.Camber on the Static Longitudinal,Lateral, snd Directional Charac-teristics of a 45 SweptbackWing at Mach Numbers Up to 0.96. NACARM A56D02, 1956. F3. Winter, K. G., Scott-Wilson, J. B., snd Davi

46、es, F. V.: Methods ofDetermination and of Fixing Boundary-Lqyer Transition on Wind Tunnel *Models at Supersonic Speeds. R.A.E. TNAero. 2341, (British)j1954.(Also available as A.R.C. rep. 17,416, C.P. 212, 1955, and AGARDrep. AG 17/p7, 1954, pp. 167-191)4. Main-Smith, J. D.: Chemical Solids as Diffus

47、ible Coating Films forVisual.Indications of Boundary-Layer Transition in Air and Water.R. :% 62.223 .* pg 2.304;:%1:?.% 2.132 1.U83.340 66.9036.623 1.931 3.340 1.393 2.13265:9031.5* 1.93171.452 l. 6.623 1.9.850 1.8.452 l.1 .872 1.468 9.550 1.93313.0231.953 m.qo 1.217 75.572 yg13.023 2.155 80.17019.2

48、13 2.194 84.352. .953 19.ZL3 2.317 8k.3523.SXI 2.333 88.421 :963.m5 S.2C0 2.382 88.42130.%7 2.4W 92.I!flk .473 .P536.6M g.m 30.997 2.429 92.384 .4732.5H42.oY) .238 36.610 2.5U *.212 .2382.541 J 1OO.(MO .m 42.c50 2.541 100.OOO .WI+e mum: 0.190percent chord: 0.236pcrccntclmrdGZz z:%eadlng+igcradius:0.

49、370percentchxd0.5E0pcreentchordLeadfag-edgeradius:0.713percentchord%eaWg-edgeradius: 0.924pcrccntchordProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-X2 NACTN 4366TART.T! 17. . TTiTIWY llAWA TiVfiTTIXi!.Q- +. .,MU w - . ”.LFigure Variables M - 6 5 Transitionwire3 cm) ag. CL 0.25 to 0