NASA NACA-MR-L5L11-1946 Two-dimensional wind-tunnel investigation of two NACA low-drag airfoil sections equipped with slotted flaps and a plain NACA low-drag airfoil section for XF.pdf

1、MR No. L5LU % NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ORIGINALLY ISSUED January 1946 aa Memorandum Report L5Lu TWO-DIMETlSIONAL m- IXVTSTIGAIION OF IUO WA K)W-DRAG AIRFOIL SECTIOIUS EQUIPPED WITTI SILTED FLAPS AND A PzAIlV NACA =DRAG AIKFoa SECTION FOR -1 AlRPUNE By Iamnce K. Loftin, Jr., and Fr

2、ed J. Hce, Jr. liaflleTr r two- di:nensional, low-turaulence tunnels of three 24-lnch- chord airfoil models yepresentin; the root acd tp airfoil sections and an intermedfate nirfoLl section of the p?oposed- Chance-Vought x6v-1 airplane. tested were the 3ACA 65(215 )-114 (root sectfor,) , the TJACX 6

3、51-212, a = 0.6 (tip section), md an intsrmediate section taken at approximately 55 percent of the semi- span. were equip;?ed with slotted f laTs. The airfoils The models of %e root and intermediate scctions The tests included the determination of tile aero- dynaxfc characteristics of the t?Jee slai

4、n alrfoil sections in 30th ths smooth cond;tion and xitk standard roughness applied to L1-6 leadin2 scige. Lift tests of the root and int6rmsdlate airfoil sections werz made for a ra9i;c of flap dcflections extending Lrom 00 to 50. ?lost of the data werd obtained at a Reynolds nunber of 3 X 10 altho

5、ugh cornparison tests wgre conducted at numbers of 1 X 106, 3 x IO”, and 6 x 106. free - str e an dynax-ic pr e s sure airfoil section lift aSr f oi 1 se et ion dr a6 airfoil sectj.or, qu-arter-chord pitching rioliient airfoil section lift coefficient, L qc Zmax maximu% airfoil section lift coeffici

6、snt, - tic d airfoil section drag coefficient, - qc Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-m NO. 511 3 air f oi 1 s e c t ion .duar t er -chord pit chinG-nome nt CmC/4 ccefficient, ma I vL airfpE 1 se c t i on pit c3Lq -mmcnt c oef fi cieiit

7、 ma.c. a5out the aerod-JnaxBc cm.Ler, Cm,.C. a airfoil. section angle of z,ttack 6 flap deflection with res1:ect to airfoil chord a airfoil section r-,eynolds Eu:nber The airfoil sections for which data wei-e desired co:isisted of the NACA 65(215)-1114 (root section), the -:AAX 652-212, a = G.6 (tip

8、 section), agd a:i intermediate airfoil section taken at appoximacely 55 peicent of the sexispan. The Foot and intermediate airfoil sections were eyiiip2ed vith slotted flaps of 25.92- arid 33.62- 2ercs:t airfoil c?-crd, ri .-. s.,ct:i .v.ly. I.:;:,(-; ;-:.3:;.lt111.- cl;crd sizes correspond to a fl

9、ap of const.ant chord length on the three-dimensional wing. With id intermediate airfoil sections, respectively. I sections tested xere constructed of laminated mahogany. The mrfaces v:ere then painted and sanded with number n.llar to that desci7ibed for the plain airfoil models. Drawings of the roo

10、t and interaediats airfoils sectiGns wf-t;h flag deflected are presented in fiprss 1 and 2 i;o$,etiL.!Cr xit2-i the flap ordinates and the dimensions locating -tnat.ed by the Chance-Vought Cor3oration. A dyawing of the tip section is shom in. ffpm 3, The 2!+-fnch-chord models of the three airfoil Pr

11、ovided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-T;1;SllS The tests of the tlwee airfoil models were con.;ilc5 i:ieas-re 3 feet by 7.5 feet; the models, when mounted, coinpletely span the 3-foot dfmension wNtt.h. the junction between the nodel xnd t,Li:

12、mti walls sealed., gratiqg t. I3iqessupe reaction 0i3. the CCT of the turnel. Drag coeflicisnts werz determlnsd by the wake-swve-2: method, and the quarter -chord pitchi1;iz-n?oment coefficients vzere measured with a torL:ue balance A11 coefficients :./ere calculated usinc t9.c basic airfoil chord i

13、th flap retracted an6 neutrzl a A !;.,ore co?.cplEtt,e descri2tion of these tunzels and the ixetliods emFloyed for obtainin? and Fedticinz the e:.iperIlriental data 4s contailled in reference 1. from refereqce 1 xere used to correct the tmmsl data to f r e e - af r c ond i t i on s : 9 JT t VJ 0 - d

14、i ;.ne 11 s 1 Lift qeasu.rencnni;s wers. o9tained by inte- ar!.d Cei.Iil*lg The folloWin fcrn;ilas derived i - - l.G594q! a, := l.Olcja, Lift and drag res-Jlts were obtained for the three plain airafoil sections in tl-16 snooti7 condlt Loil at Keyrzolds numbers cf 1 ,Y 106, 3 x 190, 6 x loG 2nd 7 K

15、lo6. and dr-y were also yeas:ired at a ,ie;nolds nu-nber of 6 the leading edge of the ?.rodel. Tlie i)itc?Anz-nolmnt cha-acteristics of t!ie tzee _iodsls i-7 t-: s.;oot- dition vrere CleteriiJred at Re;-nolds n:m?ers of _I :. 10 , 6 x 106, and / Lift 10g“witb standard rouepness (reference 1) ap:Jlie

16、d to LL “2- - x 106. Lift results -ere obtained iw t:TLE: roct aril inter- niediate airfoil sections i: tbe slioctl- conCi tion thro;gki, a ran:“ of f Lap def lecti 3ns ext.nd: nr b fro.:? Oo to 50 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NR N

17、O, 511 5 The lift characteristics of these two airfoils were also determined with the flap fully extended but not deflected, and with the flap partially extended and deflected bo. The latter configuration was intended to be used as a cruisbg deflection f0.r the full-scale airplane. The Reynolds nimb

18、er of most of these tests was 9 x 106; how- ever, the effect of increasing the Reynolds nxdxr from 1 X lo6 to 9 x lob was determined for the airfoils. with bo0 flap deflection, Oc full extended configuraticn, and 4-O cruise condltion. were also made at a Reyzzolds iiwmber of 6 x 10 6 . with the flap

19、 deflected 40 and standard roughness aplied to the leadine; edge. Pit ching-mome nt char act eri sL_i c s were de t erdned for both-models at flap deflections of tCo a d hOo. The Reycolds niimber of the tests was 6 x 10% xitb the flap deflected LO0 and 9 x lo6 for the bo deflection. BraC results wer

20、a obtained for oly one flap deflec- ticn, the bo (partially extended) cruise confizurati n. The data were obtained at Reynolds nixll3er-s of 1 X 108 and . 1 _. I. 9 x 106. RZFLTS AZTD DISCUSSIOIT Flap retracted,- The results of tests of the three plain airfoil sectiolzs am presented irk fig;ul“es b,

21、 5, and 6. A Cornparison of $hess results indicdtos that at a Reynolds number of 9 x 106 all three sections have agproxirnately the same inaimurn lift coeff icie;it Decreasing the Bepolds number from 9 106 to 3 x lo6 apixars to cause a decrc- ment in maximum lift coefficient of abofit 0.1 for the in

22、termedihte and tip sections, an6 0.05 for the root tion. A further decrease in Reynolds nuxber from 3 X 10 tc 1 x 106 results in a decrement of approximately 0.35 in t“le maximma lift cosfficiont for all three sections. I+, Is intsrestinr- c. to :_ate tliat tiit, maxi.r,iu.m lift cosf- ficients of k

23、he. s?!onth sections at a 3agnolds nu:ibor of 1 x -106 is of t;ie, sage ordcr of m.nitue as tkoso obtained at a 3epolds nuEbsr of 6 x log with standard roughness apalied to ,:?e airfoil leading edge. .The minimux drag coefficient of t:?e three smooth sections is seen to be approxinately 0.004.0 at a

24、 Reynolds “z Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-LZ XO. 511 6 num3er of 9 x 10 6 . 9 -77 an f*?erSfi:=nt of anproxinately 0.0010 for the root and internediate sen,tioxs and 0.0005 for tL;a tip section. The nost adverse effect of decFeasi2

25、g Reynolds number is, however evicent outside the ranp 5f lift ccafffcfents corLesnond;ng to low-drag, DecreasinL the 3eynolds n;mber to 1 X 106 resii.lts ii1 a fgrthzr increase in dl-sg, ivith the xost pronounce deflection are prcse ted / x 106, 6 x 106, and 3 x 10 , and for A Xeynolds number of 6

26、x 106 with standard roqkncss ai3:;lied to the airfoil leading edp. Included also in these figures are the data obtained at lieynolds nu;Ymrs of 1 x 106 and 9 x 106 for the (partially extended) flap deflect: 1 on and the 00 (full extended) flap conficuratioz. The maxi-mm lift coefficient obtained qrr

27、it:i the 40 flap deflection ap?ears e Reynolds number is decreased from to suffer little 9 x 106 to 3 x 10 number from 3 x 106 to 1 x 10 results in a decrexent in maximJm lift coefficient of approximately 0.32 for the intermediate section and 0.47 for tke root section. The decrement ii1 maxinun lift

28、 coefficient on botr_ plain air- foils for a decrease in Xeynolds nuxber from 3 X 106 to 1 X 106 was of the order of 0.35. As has been noted in tests of an Girfoil e;l+ippeiZ with a doilb*le-slotted flap (refeyence 2), the anzle of zero lift decreases with decreasing Zepolds nurn9er. in maximum lift

29、 coef ficient res:i.lting from standard leadingedge rouC:mess are 0.06 xore and 0.13 less, respectively, for the airfoils wrtli LO0 flap deflection than foy the plaln airfoil. This result seems to agree with the data presented in reference 2 which shows that for an. airroil apyroxiaately 14-perccnt

30、thick e iL:i-ped with a double-slotted flap the decrement in !11;3iirmi lift coefficient resulting as a cmseyuennce of standard roughness is approximately the saxe for ti2 plain airfoil as for the airfoil with fla? deflectcd, As was noted ?nJitk: the pla.n airfoils, the maxirnux lift coefficient of

31、the smootk sections at a Eegr,olds .number of 1 x 1.0 are nearly the s me as ti:ose obtained at a EZepoldv in figures 10 a?d 11 *Cor :eylds nuabors of 1 x 1b 8 , 2 ts t ho-:Jever, ecreasing the Reynolds For the root and lnterniediate sections, the decrements 8 number of 6 x 10 z jvith standard leadi

32、ng-edge roughness. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-D a; data corresDonding to Reynolds numbers of 1 X 10 t;? and 9 X 10 6 are presented in figure 12 for the root and Intervediate airfnll sectto34 vith the Lo (uartinlly extended) flap

33、deflection. Coynparisan of these data vith those obtained for the plain airfoil shows that the flap causes an increment 5.n minimum drag coefficie t of amroximstely 0,001.0 at a Reynolds iiumber of 3 x log on both sections. At this sanie Reynolds number, tkie use of the 4.O flap defl.ection caussd t

34、he lift coefficient correspond! nE; tg the uner 3.imit of the low- dras rar-e to imreese rroiqi 0.3 to q.L!. for tlie root. section and fro:? 0.35 to 0.5 fo;? t:e iaterriedicte section. The pitching-riornent characteristics of the two air- foils with flap deflections of 4. (partially extended) and L

35、Oo are presented 5-11 fisure 13. A cor,y3arison of these data with those in reference 2 for a 6-series air- foil equi?:xL: with a double-slotted flap infiicates that for lift coefficients up to the stall the pitching- moment coefficients are considerably less for the air- foil with the slotted flap,

36、 C ORC LUSI OITS The results Qf a two-dimensional wind-tunnel inves- tigation of two NACA 65-series airfoil sections of approximately 14- and 13-percent thickness and equipped with 25.92-nercent airfoi 1 chord and 3 3.62-percent air- foil chord slotted flaps, respectively, z.nd a ?lain airfoil secti

37、on indicate the following conclusions: lift coefffcient was 40 for both of the airfoil sections equin?ed vith slotted flaps. 1. The orJtimum flap cleflestion for maximm section Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-9 mi NO 511 3, At a Reyno

38、lds number of 9 X 10 6 the use of a 4O cruising flap deflection caused the lift coefficient corresponding to the upper limit of the low-drag range to increase from Oe3 to 0.L. for the root section and fron 0.35 to 0.5 for the intermediate section, On both afrfoils the Increment in minimum drag coeff

39、icient caused by the 4.O flao deflectton vas apprgximately 0,001 at a Reynolds number of 9.0 x 10 6 . Langle y Ikmori a1 Aer oneut 1. cal Lab or at or y National Advisory Comitteo for Aeronautics Laneley Field, Va. 1, kbbott, Ira B. v3n Doenhoff, Albert E. and Stivers, Louis S., Jr.? Summary of Airf

40、oil Data. KACA kCFl No, L5C05, 19r5. 2. test, TM 914. _ Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MR NO. L5Lll Figure 8.- Lift charscterletice at various flap deflections of the NACA 65(215)-114 sirfoil (intermediate section) equipped with a sl

41、otted flap for the Chance-Vought Xi?6U-1 airplane. test, TDT 915. R = 9 x lo6; Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MR NO, L5Lll Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-flap deflect

42、ions of the NACA 65(21 )-lib airloil (root section) equipped with a slotted flap for tze Chance-Vought XF6U-1 airplan Tests, IDT 915 and LTT 418. MR NO, L5Lll Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MR No. L5Lll Provided by IHSNot for ResaleN

43、o reproduction or networking permitted without license from IHS-,-,-MR No. L5Lll Figure 1 .- Dra characteristics of two configurations of the NACA 2 5(2151-fl.4 airfoil (root section) and an intexmedlate ahfoilsection, both equipped with a slotted flap, for tkc Chance-Vought XFU-1 airplane. Tests, a

44、)T 94 and 915, and LTT 417 and 418. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MR No. L5Lll h rl rl ai d 4J k ai a a Ln ai d rl cu h d L v ax mD 08 A rl rl ai d I. k ai a 4 d ki a G- aim- Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-