NASA NACA-TR-942-1949 Investigation in the Langley 19-foot pressure tunnel of two wings of NACA 65-210 and 64-210 airfoil sections with various type flaps《带有多种类型襟翼的NACA 65-210和64-2.pdf

《NASA NACA-TR-942-1949 Investigation in the Langley 19-foot pressure tunnel of two wings of NACA 65-210 and 64-210 airfoil sections with various type flaps《带有多种类型襟翼的NACA 65-210和64-2.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-TR-942-1949 Investigation in the Langley 19-foot pressure tunnel of two wings of NACA 65-210 and 64-210 airfoil sections with various type flaps《带有多种类型襟翼的NACA 65-210和64-2.pdf（23页珍藏版）》请在麦多课文档分享上搜索。

1、REPORT 942INVESTIGATION IN THE LANGLEY 19-FOOT PRESSURE TUNNEL OF TWO WINGS OFNACA 65-210 AND 64-210 AIRFOIL SECTIONS WITH VARIOUS TYPE FLAPSBy JAMESC. SITELM and STAPUJZTH. SPOONEESUMMARYAn investigation has been conducted in the Lungky 19#ootpremme tunne to determine the mom-mum li and stallingcha

2、racten”stics of two thin un”ngsequipped with sereral types of$aps. Splil, single s.?dted, and double slotied$aps were te8tedon one u+ng which had IV.ACA 65I?1Oairfoi sections and splitand double 8fotted $aps were tested on the other, which hudNAG!. 64fi10 airfoil sections. Both m“ng8 had zero 8xeep,

3、an aspect ratio of 9, and a taper ratio of O_J.+lt a Reynolds number of 4,400,000 each type of j%p in-creased the mm-mum lift coej%ient8 of the two ving by incre-ments which were appron”mately proportional ti thejlap neutralmlues of 121 and 135 for the lITACA 652?10 w.ng and theAJ.AC.464+z?1O u-ring

4、, respectitvly. l%e caues of mam.mumLijl coefim”entfor the unkgs with full-span double slotted jlap8were 2?.48 and 2?.76, which ralues represent increment8 of 105pcrcnt oj the$ap neutral ralue8. me addition of a repr.%enta-tice fu8elage or leading-edge roughne88 was more detrimentalto the NXA 64g10

5、mung, but its ralue8 of mm-mum licoej?cient were still consistently higher than those of the .Y.K.465+?1O wing. The rakes of mazimum lifi coejitient increasedwith increaa”ng Reyno.Hs numbers up to a ralue of 4,400,000.Abore this ralue, the test Mach number was high enough 80 thatthe eects of compres

6、sibility appeared b cause the ralue8 ofmazhnum lifi coefficient to increase le8s rapidly or to decreaseunlh increasing Reynolds numbers.The Wall of the .dp. 64-g10 wing wa8 somewhat moreabrupt but slightly farther inboard than that of the ALMA66-210 ming. Tie pattern of the 8tai was approximately th

7、e8ame for all $ap con$gurations with or without leading-edgeroughness. The main e$ect of roughness was to make the stulprogression more gradual. The fuseage, houxrer, caused thestall to begin inboard near the un”ng-fuselagejunction.INTRODUCTIONThe wing sections of an airpIane capabIe of flying at hi

8、ghsubsonic speeda must be relatively thin in order to deIay theonset of the effects of compressibility. These thin sections,however, cannot normally develop as high values of maximumlift coefficient as thicker sections used on sIower airpIanes.More powerful high lift flaps must therefore be used on

9、high-speed airplanes to obtain landing characteristics approachingthose of Iower-speed, but otherwise comparable, airpIanea.In order to develop high lift flaps suitable for thin airfoils, aninvestigation was conducted in the Langley two-dimensiomdlow-turbulence tunnek. (See references 1 and 2.? The

10、mostpromising results of this irmestigation were incorporated inthe design of two thin wings, the t.hree-dimemional charac-teristics of which were instigated in the Langley 19-footpresmre tunneI.One of these wings had NACA 65210 airfoil sections andwas equipped with spIit, single slotted, and double

11、 sIottedflaps. The other wing had N.*CA 64-210 airfoiI sectionsand was equipped with split and double sIotted flaps. Theplan form of both wings -wastypical of a Iong-rsnge airplanein that the aspect ratio vras 9 and the taper ratio was 0.4.Presented herein are the results of tests made at relatively

12、high Reynolds numbers to determine the maximum Iift andst.ahg characteristics of these two wings tit h partiaI-spanand fall-span flaps both with and viithout a representatiefusdage and Ieading-edge roughnma.COEFFICIENTS AND SYMBOLSThe coefficients and symbok used herein are defied asfouows :c. lift

13、coefficient (L/)CD drag oefficeut(/cm pitcl$g-moment coefficient (M/i7)cLfi,=L+= (Tail length=3E)c. mazACLWwhereLD.-1!z.SEPT“T“,cbYand:MPamaximum Iift coefficientincrement.in CL_ due to flapsliftdragpitching moment about 0.25Zdynamic pressure of free stream ();P1.”2wing area (24.94 ft.mean aerodynam

14、ic chord (1.769 ft) (fcdy)mass density of airairspeed-rertical velocity in glideIocal wing chordwing span (15 ft)spanwise coordinatecorrected tmgIeof attack of root chordReynoMa number (PIZ/p)Mach number (-/a)coe%icient of ticositysonic velocity419Provided by IHS Not for ResaleNo reproduction or net

15、working permitted without license from IHS-,-,-420 REPORT 942NATIONAL ADVISORY COMMWE FOR AERONAUTICSMODELS AND TESTSThe two wings were constructed of soIid steel and weregeometrically similar except that one was contoured toNACA 65-210 airfoil sectionsand the other to NACA 64-210airfoil sections, T

16、he taper ratio was 0.4 and the aspectratio was 9. The sweep and dihedral at the 0.25-chord linewere 00 and 30, respectively. Both wings were uniformlytwisted about the 0.2 mp?ct mtlo, 9,01;weahont, ; bper ratio, 0.4. (All dtmeoshms are In tnohm.)Provided by IHSNot for ResaleNo reproduction or networ

17、king permitted without license from IHS-,-,-IN YESTIGATION OF TWO WINGS OF NACA 65-210 AND/- -Wq NACA wnO w.(CJ DonbIe slo+ted flaP; N-ACA 65-!210 fig. (d) Double slotted flnp; XACA 3fs 0.17.FLAP EFFECTIVENESSIf the values of maximum lift eoeflicieritof the wings withflaps aro expressed in percent o

18、f the flap neutral values, thaflap effectiveness for both wings was practically the same ata Reynolds number of 4,400,000. Inasmuch as the flapneutral value for the NACA 64-210 wing was 1.35 as com-pared with 1.21 for the NACA 65-210 wing, the flap ex-tended values for the NACA 64-210 wing were cons

19、istentlyhigher. The increment in mtMnum lift coefficient due toflaps for the smooth-wing condition and for a Wynoldsnumber of 4,400,000 are as folIows:1-Putfd . . . . . . . .shut . . . . . . . . . . . . . . _ u- 0:ffSlotted . . . . . . . . . . . $:; :-: :HDoubIedotted.- . . . . . Full-: :# o.b2 35t0

20、3a,59 44. ds.1.07 74to%L 41 10sProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-INVESTIGATION OF TWO WINGS OF NACA 65-210 AND 64-210 AIRFOIL SECTIONS WITH T“ARIOUS TYPE FLAPS 423a cm c-FIGCBE i5-Aemdynamfc chwscteristks of NACA 6+210 wing with snd wft

21、l.umtstngIe slotted P M=$ 0.17.The results of the tests of the NT.ACA65210 wing with theflaps removed from that part of the wing normally occupiedby the fuselage are shown in figures 4 to 6. The data in theIiuear lift-curve range indicate that some of the lift due tothe single slotted and double slo

22、tted flaps was carried acrossthe fusdage, whereas practically none of the Iift due to thesplit flaps was carried across.For all configurations the fusehge caused a destabikingeffect on the pitch moment equal to a forward shift of theaerodmic center of about. 5 percent of the mean aero-dynamic chord.

23、EFFECT OF LEADIXG-EDGE ROUGHIWSSLeading-edge roughness caused a rounding of the lift-curvepeaks and a reduction in the maximum lift coefficients ofboth wings with and without flaps. The reduction USudlyamounted to about 0.2 for the h7ACA 65210 wing and about0.3 for the NTACA64-210 wing. As was true

24、for the fuseIageconfiguration, the maximum lift coefficients of the N7ACA64210 wing were higher than those of the NACA 65-210wing even though t-he effect of roughness on the NACA64210 wing was greater.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4

25、26 REPORT 94 .f 0.17.At low angles of attack, the addition of leading-edgeroughmm usually decreased the lift coefficient sIightly. Forthe NACA 64-210 wing with double slotted flaps (3fss0.17,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-INJESTIGATI

26、OX OF TWO VKKNGSOF NACA 65210 AND 64210 AIRFOIL SECTIONS WITH VARIOUS TYPE FLAPS 429FIGURE1L-“o + -.2 -3 =-.4 7mdynamic 2.0d3/.6 /-Q.4 163 56 7xl 34 5xfo6Raymb% tips neutral. R= 4,400,MW;M- 0.17.1.8M1.4q L2/.0.8“54 O 4a8 12 MRough fiOWMet-m”+tentstalla-S.2 q “ /.55-.- ._.u- /0.3” CL-L63a-107” CL-1.5

27、5L81.6!4q 12t.o.8“G4 O 4a8 12 /6Ccmpkte stoflElJ, Jcm.ss flow/.8t.6/.4q 12Lo.8“54 O 4a8 !2 16-wa-8.F CL-142 a-Z5” c “ L32a-9.J” c - L46 a-86” w CL-1.41a=i02 CL-I.47 a=9. T w c-147a=tLr CL- L37Smooth WFmuEE 19.-Stdl oharactwfstka of NACA OHIO wing; lalg spilt flqm i?= 4,4C0,0X3 .M= 0.17.a=/L6” CL-L44

28、 a=l CL“ L4iRough kadi edge Smoih wing md fuselageProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-INVESTIGATION OF TWO WINGS OF NACA 65-210 AND 64-210 AIRFOIL SECTIONS Wl?lTH VABIOUS TYPE FLAPS #aRo ?%W.Ma-nn”ttent Sfdfa- 7.? CL- 68G”m “.u2LoL6/.4/2L

29、o.8.6-4 0 4e8 12 f6Cwete Staaa75J, J0-0ss fbwa- 73 CL=/64a= full+pan sp13rRam. R= 4J131Jco;.Wr=0.17.e= 89= =L86 a= z? CL=i a= 6t. 1J, /.4i. Cross flow /, 048/216 “- /“Q4o 4juJ-LLJ6 ,hermitten t stall $4$:4zgp#la a aMmE!%l-20 w 20 M m1.8 Rough ffOW L8 j-6 G 1.6t.4 14 IJ, I L41VHH1?a= ZY lq = 203 a= t

30、19 =im a- 640 CL“183a- 85 CL 200 a- 7.4* =lfi? a- tL3” CL=77a=S4* CL=/.80 a“ 88” CL=L-78 a- Q3 -L67Shxmfh Wih$l Rough kgoiq edge tioth whg md ikebgeFIQUM !23.-StaU ekmcterfatlcs ofI?ACA 13E-210w; tlal.apdonble dotted flap% R= 4#M,fK0; Mu 0.17.Provided by IHSNot for ResaleNo reproduction or networkin

31、g permitted without license from IHS-,-,-INVESTIGATION OF TWO WTSGS OF NACA 65-210 AND 64210 AIRFOIL SECTIONS WITH VABIOUS !ITPE I?IAPS 437a75Ccu7ete sfdla7511cross m.e =.26- Q Ms= Eli.1.4!2McL #.6.4404812f6Q-=14.Cr cLi25a+5. o“ CL=/-32 tY”14_3” CL104”Snaofh uc?qct=14.w CL“f.wRough Iecm%gecsL4E42;.6

32、.44 O 48 12 i6ffi24 CL=Lf3wcY=f2.9” c! =Lfi.c=/4.4 CL=L/8-Q-14.SMOi% wig and fuselcgeFIGtJEE 2/.2 1/, LMermi?tent stall 0 Cress flow L4048/216 .w et -cf-tL4” CL=L70 cr”w “ - c, 4.3f cY=t14e CL/.40-w=12.4” CL=.178 VL?6 cr=9.2” CL=l.48.u=12.9” CL=L82 u “9. 7 cL fi a=/O. 7“ CL-(%)4r/9.2” CI=L52 Qf=l pa

33、rt!e.kpan spIIt fMs 0.17.2!01.8/.6 f.41.21.0284 0 4a8 1216H./?FKGUM 27.-St8ULuSaharwtehtiaofNACIA W21O W fulI spIIt flaIxI. R u 4,4CKLm M= 0.17.Rough Jeodingedge %?oofh whg omd fuseicqeProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-INVESTIGATION OF

34、TWO WINGS OF NACA 65-.210 AND 64210 AIRFOIL SECTTONS WITH VARIOUS TYPE FLAPS 4392.4222?0q laJ!6L4l!2-4 0 4U8 L? i6-a= 8.7 =230 a= % $4491a. ooa9:894 3.65614.SW 4.233la 909 4.%24.921 h207.22.a36 6.72234.931 6. e6439. 2U9 O.m44.6s4Mt.m k%65.014 6.625a. .217% 4.71270.043 4. lm7L 045 3. 4)Q8LC44 .zmS5.O

35、aa z M790.023 L 32795.014lm. Ceo :%Lmrcr nufweSt,atlono.665i EZ692h 1021;%15,101m WI26.07930.0043K 0-4940. n3246.018M%lw.97304.96469.95774. Q5379.950M. 902W. Q72M ml100. mlOrdfnntoo-.710-.839-1.069-1. Z?S-1.839-2. ax-4.621-1 ma-3.340-z. 60747S9-3. SW4. a2d-3.% :99Zal .mNi -.70+“ .al114).17.I I SmoOt

36、h wingmaptype I Fp sn mFlrlps nentral-. - L21rplIt ._- . . . . - oL- :. -Slotted - =:-;: srtiaL - !2.10Double st F-2.4sJ.=IT-*L36 L20 L?riL3i 1.57 LSIL% 1.6S LSi.- 1.i5 -,._-_ Z.) -L42 L94 226276 22i 2.5515-21064-210 5s-210 64-2101111mamw1.12 1.17 rL99 1.041.49 L5S L49 L=1.2-t LS6 LiO L75LZ -. LX -101 - LW -2.01 2.00 LE8 2.022.34 9.56 2.s 2.4sW-47795 129Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-