NASA NACA-RM-A54J12-1955 Pressure distributions on triangular and rectangular wings to high angles of attack - Mach numbers 2 46 and 3 36《当马赫数达为2 46和3 36时 三角形和矩形机翼大攻角上的压力分布》.pdf

1、1. - -.RESEARCH MEMORAND uPRESSURE DISTRIBUTIONS ON TRIANGULAR AND RECTANGUiXRWINGS TO HIGH ANGLES OF ATTACK -MACH NUMBERS 2.46 AND 3.36By George E. KaattariAmes Aeronautical LaboratoryMoffett Field, Calif.”. . . . . . . . . . . . . . . . .lt!icnmtdml contains InformationLKeQtlugtinwad mfOIMeof Cali

2、bration of the air stream indicatedthat the value of( - Po)/ at M = 2.4-6was esstially 0, but that at M = 3.36 it. was approximately0.01.Chordwise pressure distributionswere integrated for each spanstationby a tabular method to give local span loading coefficientCCn and local center of pressure z/c.

3、 The absence of orfiices at theleading and trailing edges of the wings required extrapolationsof thepressure distributionto these points. Linear extrapolationswere used,based, respectively, on the pressures measured at the first two and lasttwo orifices of each span station. The spanwise load distri

4、butionsweresimilarly integratedto give total load CN and center-of-pressureloca-tion /cr and /s. The span loadingsbeyond the most outbosrd stationof the models were approximatedby assuming a parabolic load distributiontangent to the slope passing through the loading of the last two out.board station

5、sand falling to zero at the tip.Validity of DataThe validity of the data is affectedly measuring accuracy and toan undeterminedextent, at the highest angles of attack,by plate-boundary-layer interference. The slightvariations from constant testconditionsand inaccuraciesin setting the model angle of

6、attack causeda probable error of less than iO.02 in the pressure coefficientsat bothMach numbers. The effect of the boundary-layerplate on the semispanmodels was discussed in reference 4 wherein it was noted that the root-chord pressure distributionof the unthickened-rootrectangularwingProvided by I

7、HSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 NACA RM A54J12comparedwell with those predictedby shock-expansiontheory at Mach #numbers 1.45 and 1.97. Good agreement indicated that the boundary-layerplate had little effect at the root chord b-elowthe angle of

8、shock-.detachment. The pressure distributionat the most inboard spanwise.station y/s = 0.025 was also in good agreementwith theorybelow theangle of shock detachmentfor Mach numbers 2.46 and 3.36. The onlyconsistentindicationof.boundary-layer-plateeffectswas evident in thecase of the aspect-ratio-4tr

9、iangularwing when tested at Mach number .2.46 for angles of attack above 25 . A reductionof about eight percentin the span loading at the root chord occurredwhen the Re”olds number .-was reduced from 0.44x108per inch to 0.26XL08per inch. It is not clear why the other plan forms do not show correspon

10、dingReynolds number effectsatothe root chord. “Theaccuracy of the dataofor angles of attack aboveM , and those for wing 2 at angles above 25 at Reynolds number 0.26x108 -per inch, are subject to some uncertainty.RESULTSTabulations of pressure coefficientsare presented for the modelsat M = 2.46 for R

11、 = 0.44x108 per inch and at M = 3.36 for R = 0.85x108 .-per inch in tables I(a) to I(j). The contributionsto the loading andto center of pressure for each spanwise stationare presented in tablesII(a) to II(j) for both upper and lower wing surfaces.Summarized in *tables II for each wing are also the

12、normal-forcecoefficients,thecenter-of-pressurelocations,and moment coefficientsabout the wingcentroid of area. Figures 2 to 6 present plots of span loading coeffi-cients,normal-forcecoefficients,and the center-of-pressurepositions .for each wing. Data taken at R = 0.26oe-per inch at M = 2.46 tieshow

13、n on these plots for comparison. Plotted on part (b) of figures .2 to 6 are also the values for the normal-force coefficientsas predictedby linear theory.DISCUSSIONAngle-of-AttackEffectsIt was noted in reference 4 that all five wings tested at Machnumbers 1.45and 1.97 tended toward a uniform loading

14、with increasingangle of attack. This was also found to be the case for the loadingson the same wings at the higher Mach numbers of tie.present test upto the angle of attack of 40. However, on all wings tested beyond 40,the pressures on the root chord decreased soiiewhatwith a consequentmovement of t

15、he center-of-presswe position outward and toward thetrailing edge. This phenomenon is believed to be the result of inter-ferencebetween the bow shock and the plate boundary layer. The rec-tangularwing data are in fair accord with shock-expansiontheory inthe two-dimensi.on whereas with increasingMach

16、number, the normal-force curve tended to become concave,resulting inhigh= slopes at high angles of attack.7No large effect of Mach number on the center-of-pressurepositionwas noted. For the triangularwing of aspect ratio 2, in the moderateangle-of-attackrange of 3 to 25, the center-of-pressurepositi

17、onmoved slightly forward (0.03cr)with increasingMach number while above25 there was no consistentMach number effect. In the case of therectangularwing and of the aspect-ratio-4triangularwing, the predomi. nant effect of increasingMach number was to decrease the spanwisevari-ation with angle of attac

18、k of the center-of-pressureposition.Effects of ThickenedRootIn reference 4, it was noted that at M = 1.45 the span loading wasnot affected by the thickened root for either wing. The center-of-pressure position of the rectangularwing moved O.Olcr forward due tothe presence of the thickened root secti

19、onwhile that of the triangularwing was unaffected. At M = 1.97 the root-chord loadings of both wingswere reduced by the presence of the thickened root so that the totalnormal force was reduced by 5 percent in the lower range of angles ofattack (3 to 17.5) and by less than 2 percent above 17.7. The c

20、enter-of-pressureposition of the rectangularwing was again moved O.Olcrforward while that of the triangulsx wing was unaffected by the presenceof the thickened root section.The effect of thickening the root-chord section at the higher Machnumbers of the present test can be seen by comparing figures

21、2 and 5 forthe aspect-ratio-2triangularwings and figures k and 6 for the rectangu-larwings. At M= 2.46, the span loading of the rectangularwing wasnegligibly affected by the thickened root chord up to 30 angle of attack.Above 30 the umthickened-rootwing had unexpectedlyhigher chord loadingat the tip

22、, giving total normal forces 3 to 4 percent higher than thoseof the thickene whereas with increasingMach number, the normal-force curveslope tended to become concave, resulting in higher slopes at high anglesof attack.Ames Aeronautical LaboratoryNationalAdvisory Committee for AeronauticsMoffett Fiel

23、d, Calif., Oct. 32,REFERENCES1. Moskowitz, Barry, and I3d.2.40;R.O.44W perInch%2%.rm.873.949.37Y.7XSay.750.875.S65.V-6.2P.313.m:2.%0Y1.3.02!.2%.m.7X.875lU.CRMJ154J12.(b)W1nu1:M-3.$3:RsO.E6xl:.lak-.lm -.094lca ml lQ q6.=9 -ml 099 -.m -alla W * -. ;% z% .:%j :% .% :% :; :% :% y :3 :% ;“=.334 om -.094l

24、W. 1o1 103 c$g ce3 -.* -.05 -JJv . .* .W2 .1= .233 .347 .% .% .?a .94L1:402-Arm -.W5-.103 -.1o1 la:% -.301 -.C45 -.1* -.1o1 -.302 :% :% 13 :% :% .m.0i9 .(54 . w .1s9 .W7 S47 .393 .Tu la 1.757-.012 .016 .* .096 .179 .2eb.e .W .7b3 1.1* l.7s0 -.Qn -.W -.m W -.1o1 -mm ma -.C47 -.*9 -.034 -.o1o .m .033

25、.C?6 J63 a .m .* .746 1.19 1.*.673 -.099 -.* -.1o1 -.lm -J -.(EA -. -.070.%3 -.CE% -w-.03Y -.*3 -.m -Jxa .mk .ql .141 .ag .347 ,487 .7k 1.W 1.173-.027 -. -.ua -.= -.ti -.073 -.037 - .*3 -.=6 -.o1o .o14 .* 1.U3 au .W . Q3 .67 .* l. L? -. -.W -.3CR -.302 -Jo3 * -.c# 06 -X44 .013 .* J2L .173 .242 . *Y

26、.Mo .* .733 .023 .* I.uka -owl -.W3 -.lm -.lm -.1o1 -.OX -.023 - .C81 -.-.299 -.1o1 -.1o1 %Q * 072-.% -:g :% -m .9 .M .W9 .$ .%1 .m .939 Mg :.OJ :OSJ :aM-2.66: R.0.66xlM-9.33;E-O.8MCfw !mh1 1 1i).m 0.3Q 0.% 0.679 o-m 1.2%? l.ka .2LY .33J .$.% .683 .%7 1.M-2.46;IO.44xl%.242.367.492.617A:; -.1*-.1%-.1

27、* -X9-.195-.195-.1S$-.198-.204-.2l-.197 -.w9(f) Wing S; M=3.36; W0.85% :ti :Z; :% :x :g :?; :% :g-w :j$ %iJ :% %J :$ :% :% :M :% :% :2 :$ :% :% :% :;g ;:% ;:-.o76 -.WO.367 -.C93 :% -.OSU- .C!49 -.a?l .W1 .040 .= .l . .% . , .992 . b -.om -.076-.096-.0% -.03 : .: .0s .I,mcqll -.; :%.563 KJ -.*?-.091

28、-.097 -.075 o -.039 .COL .034 .OP .W .193 .292 .444 .619 .8461 .148 .212-.09.?+?gm :% :$ :;% :% :Tw.W :3 -.056-. -.C83-.074-.055-.046.9s3 -.097-.? -.094-.078 .033.Oxl.172.277.411-.C w% 46 ; R-o.44wf Pr khr 1 II %*,.-f-.In.35.W.$29.772.%75.90TI-.190 -.*-.191-J.-.193-Zm: :%-am -X4-.203-.2031z-.207-M-=

29、03-a-.209-.2SI TI-.ZU7 -.2M-aal -.2m-.203 -.=0-m -.2W4-.2!J.0-ss3 -?K-3.38;R-O.SEJ.357d.W3.6W. M-2. 46; W. 44wf w m?hI ,RM A54J120)Wing5;M-3.36;RM.65W0per InchUQpermufac. I kmw 9urfac*.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. , ,TABLE II.- S

30、PAN LOADDE?LIYIBION,NORMALFQRCE,AND CEWIZROF PRESSUREOF WGI .,.,*- I v., -w. I - *. Iiwmml: M*.*, m.lwrw!tiI .,“-I- -* -.t I %L-.-timl. - 1 -u-mu IEM.* R. O.4VIUFU WIMWIM4: II-3.* m.mfmr lmb, .Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-, a71 ,TA

31、BJ.JzII. - SPAN IAAD DIS!CRIBUJ!ION, NORMAL FORCE, AND CENTERi)whg 6;M-2.49; R=O.44P idOF PRESSURE OF_- ConcludedmProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM A5hJ12a71a15Provided by IHSNot for ResaleNo reproduction or networking permitted

32、 without license from IHS-,-,-, , ,.PlmfumI IMC61020 30 40 50 60Angle Of at fack, a, deg. 70 w(b) Normal force.Figure 2 Aerodynamic characteristicsof wing 1.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM A54J12-.21(c) Center.of.pressure posi

33、tion; M= 2.46. “. .68r.66.64,62.32wingSrea3ci 25 20 15 10 6=ls=-.34 40.36 .38Spanwise location, jVsCenter-of-pressureposition; M = 3.36.Figure 2 Concluded.42 44,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-J .5M = 2.k6.Icentroid ofwing areae.32 a7

34、134 .36 .38 $0 .42Spanwise location, jVs(d) Center-of-pressure position; M = 3.36.Figure 3.- Concluded.23Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-24 MikaaMkw NACA RM A54JI.2 -. . . . 3 IM=3.36o .2- .4 .8 lb o .2 .4 6 .8 10Spanwise location, y/

35、s(a) Span loading.2.0 IM=2.46/:1.5 1 /6hock detachment , / M=S.3.- /w: Lo / “ r / “ “ual ) / R Linear theor $ .5 i:Experirnenta data M0 R=44 x l6o 10 20 30 40 50 60 70Angle of attock, q deg. w(b) Normal force.Figure 4.- Aerodynamic characteristicsof wing 3.*.,Provided by IHSNot for ResaleNo reproduc

36、tion or networking permitted without license from IHS-,-,-NACA RM A54J12 25(c) Center-of-pressureposition; M = 2.46.2548 .50 .52Spanwise location, ls(d) Center-of-pressureposition; M Y 3.36. .Figure 4.- Concluded.54Provided by IHSNot for ResaleNo reproduction or networking permitted without license

37、from IHS-,-,-26 NACARM A54J .I I IhM=2.4621 I I I I I I Io .2 4 .6 .8 1.0 0 .2 4 .6 .8 10Spanwise location, y/s(a) Span loading.2.0:1.5aJ.-U.- /- / 9 /R/g 10 Kw/?, MLinear theory =Experiments data M1! R=44x IO /in.z 246El R=.26x I M=2.46.a71.?.68.66.64Icentroid of45ef( A.“.62.32 .34 36 .38 40Spanwise location, Jils(d) Center-of-pressure position; M = s. 36.Figure 7.- Concluded.a:; M = 2.Spanwise location, ls .(d) Center-of-pressure position;Figure 6.- Concluded.50 .52Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-