Document
REPORT No. 661
TESTS IN THE VARIABLE-DENSITY WIND TUNNEL OF THE N. A. C. A. 23012
AIRFOIL WITH PLAIN AND SPLIT FLAPS
By IRA H. ABBOTT and HARRY GRUEiVBERG SUMhIARY points along the span at the station 80 percent of the ohord midway between the upper and the lower sur- Section characteridic~ for uge in WI-W dem”gnare pre - faces. After the fhp had been set at the required wrded for the N. A. C. A. %7019 airfm”l with plain and deflection for each test, the gap between the flap and split $aps of 90 percent wing chord at a value of the efec- the wing was fiHedwith plaster of paris, which was then tice Reynolds Number of about 8,000,000. The j?ap painted and rubbed to produce a smooth, fair surface defections corered CL range from 60° upward to 76° down- of the proper contour.
ward for the plain jlap and from neutral to 90° dmimward The other model was used for the tests of the split for the split jlap. The split &p was aerodynamically flap. A O.ZOC split flap was made of brass for each flap superior to the plain flap in producing high maximum lift deflection tested and was fastened to the lower surface ooe~”enta and in ?wm”nglower proj?ledrag coe$m”ents of the model with screws. For flap deflections up to at high lifi coefints.
20°, the flap was made as a solid triangular prism. For INTRODUCTION flap deflections of 30° and more, the flap was made of The prevailing method of modifying the aerodjmamic two brass strips, each 1 inch by 30 inches, joined at one characteristics of airplane wings so that higher lift pair of long edges and kept apart at the proper angle coefficients can be obtained is to equip the wings with by eight triangtiar stiflenem equalIy spaced along the trailing-edge flaps. For the design of such wings, air- span. h either case, the flap trailing edge -was a foil section data at the proper values of the Reynolds sharp acute angle.
Number are needed for the wmious sections used elong Standard force tests were made of each combination the span with and without flap deflection. The purpose at a value of the e.flectiveReynolds Number of approxi- of this report is to present some additional section mately 8,000,000; the maximum lift coefficient was also characteristics for such use.
determined at an effective Reynolds Number of about The im-estimationcomprised tests of the N. A. C. A.
3,800,000. The flap settings covered a range from 60° 23012 airfoil equipped with plain and split flaps of 20 upward to 75° downward for the plain flap and from percent chord. The ranges of flap settings were -wry 0° to 90° dowmwud for the split flap. The range of comprehensive. The angle-of-attack range extended angle of attack for aU combinations e.tiended from from below zero lift to beyond maximum lift for all beIow zero lift to above maximum lift and, for the plain- conditions and was extended through negative maxi- flap combinations, extended through negative maximum mum Iift for most of the settings of the plain flap. All Liftexcept for flap deflections between 20° upward and tests were made in the AT.A. C. A. -rariabledensity the neutraI position.
tunnel at a l@h value of the Reynolds Number. Ma.si- PRECISION mum lift coefficients were also obtained for all combinat- ions at a lower value of the Reynolds h’umber.
The precision of the data obtained from force tests APPARATUS AND TESTS in the N. A. C. k -rariabledensity tunnel is discussed ! The N. A. C. A. -wwiabledensity tunnel, in which in considerable deted in references 3 and 4. It is these tests were made, is described in reference 1, and believed that the results may be applied with noumal the hT. A. C. A. 23012 airfoil section is described in engineering accuracy to free-flight conditions at the reference 2. The two aluminum-alloy models were stated values of the effective Reynolds Number. It made as dssoribed in reference 1, escept that they were should be noted, however, that the data presented anodically treated to provide hard smooth surfaces herein for the increments of mazimum lift due to the that could be more easiIy maintained during the course flap are somewhat lower than those obtained in some of the tests than the usual polished metallio surfaces. other wind tunneIs (references 5 and 6). The values The model that was used for the tests of the plain of masimum Iift coefficient contained in this report flap was provided with a bra= flap hinged at five may be somewhat conservative.
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374 REPORT NO. 661—NA’IIONMJ ADVISORY COMMITTEE FOR mRONA’UTICS
RESULTS AND DISCUSSION derived wing characteristics. The chamctaristics of the wing with flap neutral are obtained from tests of a Presentation of results.—The results are presented in plain airfoil.
&urea 1 to 9. Figures 1,4, and 5 show lift curves for Maximum-lift coefBcients.-The increment of Inaxi- the rectangular wing of aspect ratio 6 at both values of mum-lift coefficient due to the flap is plotted against the Revnolds Number. The other six fkures show the “ -..
flap deflection for both the plain and the split flaps in section characteristics us@Iy_~resented, wl&h “were figure .10. This maximum-lift ipcrement has been derived as explained in reference 4 and which may be plotted “instead of the more usual maximum Rt co- ““”” - distinguished from the wing characteristics usually efficient because it has been shown (references 4 and 5) presented and from previously used profile character- to be nearly independent of Reynolds Ntiber. The istic by the lower-case symbols. Thus cdOrepresents maximum-lift increment for the split flap increases the profile-drag coefficient for. the airfoil section cor- more rapidly with flap defection and reaches an appre- rected from the o@r pro@@ag coefficient ODO. ~y ciably higher value than that for the plain flap.
applying correction for tip effects, for variatio; of lift The maximum-lift incr~ments obtained from these along the wing span, and for ~urbtilence to correct to tests are appreciably lower than those obtfiined from 2.8 2.4 2.0 u’ %- .8 .4 &l .
Angle of afhdc, a , deg.
Fmurm L-LUt forthe N. A. C. A. W112 rectmdm Wh.lS 0ffW32t ~tfO 6 W~ O= fO@Ul Pkhl hp.
tests in the N. A. C. A. 7- by 10-foot tunnel (refermco the effective Rejnolds Number. The methods of cor- 5). German tests (reference 6) of the N. A. C. A.
rection are explained in reference 4 and the results so 23012 &foil v@th “imd without a 0.20c split flap de- corrected are intended to represent the section data in flected 60° were ma~e”-over a rrnige of Reynolds Num- the form required for application to wingdesign prob- bers. At the lower end of the scsJe range, the results lems.
agree with those obtained in the N. A. C. A. 7- by 10- Standard airfoil plots, of the form presented in refer- foot tunnel but, at the higher end, the increment of ence 7, for each flap deflection teeti are available upon matimum lift lies about midway between thut ob- request from the National Advisory Committee for tained in the N. A. C. A. variabledensity tunnel and Aeronautics.
in the 7- by 10-foot tunnel. Results obtained in the The pitching-moment coefficients cm ~a-,.,O for the N. A, C. A. variable-density tunnel for the N. A. C, A.
flapped airfoils are computed about the aerodynamic 23021 airfoil with a 0.20c split flap deflected 75° agree, center of the unflapped airfoiL Table I presents however, with results obtained for a simiiar model ig important section characteristics and also certain the N. A. C. A. 7- by 10-foot tunnel (referenco 5).
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2.0 2.224 T < Seciion Iifi coefficientc, Secfion lift coefficien~ q FIGUEE 3.—SWfon ohamcteristfc? for the N. A. 0. A. !2W12 afrkdlwfth 0.22+ Pfafm FImm 2-Sectkm chamcferktke f= the N. A. O. A. =E? akfoffwfth O.ZIC Pfdn 5D. km fkmd~.
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376 REPORT NO. 661—NATIOIWUI ADVISORY COMMITTEE FOR AERONAUTICS
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* Section Ii.& coe ffhent q FIGCBE 6.-Eation chamcterfstica for the N. A. L A. 2aM2afrfofl(inverted) with Fmms! 7.-E&tiorI characteristiea fcm the N. A. 0. A. 29012 afrfoil(fnverted) with OMcplain tlep. Smellflapdefleotkm. O.ZOe plafntip. L%rge tip dalxtkms.
378 REPORT NO. 661—NATIONAL ADVISORY COXIMITTEE FOR AERONAUTICS
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FITHPLAIN AND SPLITB’L=S 379
THE N. A. C. A. 23012 AIRFOIL The N. A. C. A. 23012 airfoil with and without the Simihrly, results obtained in the N. A. C. A. variable- flap shows a sudden large loss of lift as the angle of density tunnel for the N. A. C. A. 23009 airfoil with a attack for maximum lift is exceeded, except for the 0.20c spIit flap deflected 60° (reference 7) agree with cases where the pkin flap is deflected in such a manner as greatly to reduce the value of the maximum lift.
— h gened, the amount of Iift lost at the peak increasea as the mafium-lift increment due to the flap incre~.
Thus, the type of lift-ourve peak is usually either type A or type C (table 1), where the fluctuations of the type C peaks extend over a very narrow range of angle of attack and thus the lift-curve peak approxi- mates type A.
Drag ooefflcients.-fiofih+drag coefficients for the two combinations twted are plotted against lift co- &oient in figure 11. These pohr curves for the flap combinations are envelope curvcMof the series of pohirs obtained at the various flap-angle settings, thus giving at each lift coefficient the minimum proflledrag coefficient obtainable from the airfoil-flap combination.
The profle-drag coefficient increases much more rapidly with lift coefficient for both the plain and the split flap than for a good slotted flap, such as slotted flap 2-h reported in reference 9. IVeither flap can therefore FImBE 10.–Varfnt!ton with ilan deflectionofthe fncrementofse&kmmaxhnnmM be considered as suitable for improving take+ff as the memeienteaoeedbYOZOc PW fmdmut f@s on th N. ~ 0. A. ~0~ s~ofl.
slotted flap.
Mthough the plain ff ap has comparatively low profile drag at small deflections and low lift coefficients, the those obtained in Germany for a similsr model (refer- drag even with low defleotiona increasas more rapidly ence 6). Until more data have been obtained, the with lift coefhcient than for the split flap; the split flap reason for the inconsistency in the reeults from tests is sIightly superior to the plain flap in producing high of different airfoik in various wind tunnels must re- lift coefficients with lower proflk-drag ooefficierh main unexplained.
—-— 380 REPORT NO. 66l—NATIONAL ADVISORY ‘-’’---m ‘“- ‘ ‘n”” ‘ ‘T-n GUW-ML1”.L mlzl J! U.m /irlItuLYA U-L IUD REFERENCES Pitching-moment ooef%cients,-The pitching-moment coefficients for either flap are about equal for equal flap 1. Jacobs, EastmanN., and Abbott, Ira H.: The N. A. C. A.
deflection but are lower for the split flap for flap deffec- Variabl&Demlty Wind Tunnel. T. R. No. 416, N. A.
‘tions producing equal maximum lift coefficients. It ~ A., 1932.
2.
Jatibs, Eastman N., and Pinkerton, Robert M,: Tests h should be pointed out that the vaIues given in table I the Variable-Density }Yinrl Tunnel of Related .4irfoiIs are average values of the pitig moment and that, in Having the Maximum Camber Unusually Far Forward.
certain cases, the actual pitching moment at any lift T. R. NO. 637, N. A. C. A., 1935.
coefficient varies considerably from the average.
3. Jacobs, Eastman N., Ward, Kenneth E., and Pinkcrtou, Robert M.: The Cbaraotaristiw of 78 Related Airfoil CONCLUSIONS Sections from TestA in the Variable-Density Wind Tunnel.
T. R. No. 460, N. A. C. -4., 1933.
As applied to the N. A. C. A. 23012 airfoil section, 4.
Jacobs, Eastman N., and Sherman, A.Ibert: .AirfoiI Section “.
the split flap waa superior to the plain flap in produoing Ch&acterietics as Affected by Variationa of the Raynolds Number. T. R. No. 5S6, N. A. C. A., 1937.
high maximum lift coefllcients, iu having slightiy lower 6. ‘Wensinger, Cad J.: Wind-Tunnel Inv~tigation of Ordinary profibdrag coefficients at Iift coefficients useful in and Split Flaps on Airfcile of Different Profile. T. R.
take-off, and in having smaller pitching-moment coeffi- No. 554, N. A. C. A., 1936.
cients for equal maximum lift coefficients. Both types 6. Doetsoh, H., and Kramer, M.: S@ernatio Airfoil TasfA in were unsatisfactory in producing low profile-drag ti. Large Wind TunneI of the DVL. T. M. No. 852, N. A. C. A., 1938.
coefficients at lift coefficients useful in take-off as com- 7. Jacobs, Eastman N., Pinkerton, Robert hf., and Graonbcrg, pared with lo-w-drag slotted flaps.
Harry: Teata of Related For~-ard-&rn.ber Afrfoils in the Variable-Deneity Wind TunneI. T. R. No. 610, N. A. C. A.
1937.
8. Jacobs, Eastman N., and Rhode, R. V.: Airfoil fkctlon Characteristics as Applied to the Prediction of Air Forces and Their Distribution cn Wings. T. R. NO: 631, N. A- C.
A., 1938. “- ““ 9. lVenzinger, WI J., and Harris, Thomas A;: Wind-Tunnel LANGLEY MEUORIAL AERONAUTICAL LABORATORY, In+estiiation of an N. A. C. A. 23012 Airfoil with Various NATIONAL ADVISORY CO~M&EE FOR AERONAUTICS, Arrangements of Slotted Flaps. T. R. No. 664, N. A. C.
LANGLEY FfiLD, VA., Januaiy 21, 1938. ” A., 1939.
THE N. A. C. A. 23012 AIRFOIL WITH PLAIN AND SPLIT FLAPS TABLE I.—CHARACTERISTICS OF N. A. C. A. 23012 AIRFOIL WITH 20-PERCENT-CHORD PLAIN AND SPLIT FLAPS Derivedand eddftionslcirerecter- ~db &~mUYbe usedforstrnc- Omt Eep clareIEell[fon
——rl——l—
Wing oherecterfs
8. c. (percent c
ties A =.-round Akfoff from r/4) %uu o (per PD SE % >hord Clw= d. C-(=*J@ deg.)
r
(s2.)
d.
Ahead Above (J)
@l m I (9
W m W P) P. 9 W
--d
— ------ ----- PIain____ -00 aa 0. Sa 144 1.119 0.4
.— ---- .._do--.– -50 8.4 2.04 –15.6 . w E
------ --- .lw .--do ------ .5
------ . :~ .!: i; -:; .078 z ..-.do- . . . . . .—-. ..—--- --—----——
-–--”l”
.-—-- ---- -30 .095 .-.do ----- .2 als -------- –-—-- 417
..---- --- -30 :: i% –t ; .0S6 z ...-do...-.- .--— –. ls ------- ------ ass .-:E-
------ D2 -20 &4 L02 8.0 .342 140 .1 .16 –-.--- ------– ...-do-.._ . Oloa 407 . OIlo .. ---- D2 -15 169 0 .12 --.---– -.--— .._do. ----- .0035 417 . M185 .----- D2 -10 :; kg :; :W 0 .6s ------– -------- .._do-. -... .0072 481 . Km .-. .— D2 -5 L&? . lm % 0 .04 -------- —-—- .. ..do ------ . anl :. .0071 ..-— D2 .099 .05 .01 .---..– ---—-- .._do...-_ .0071 .0071 . . ..- :? H :2 :; .Ooa E .15 0 ----.-.- ------- .. ..do . . . . . . . Ix170 42a .0072 C12 % 0 a4 L 74 –:: . lm 249 .c& -. w 12 7 ._do----- .0270 484 c12 . ..- 0 L26 .049 m ---- .Mr3 L8 4 .-.-do . . . . . . ------ 431 --:!!!?
ao a4 L&l 46 .007 26s .25 –: OJ .:..-:: ..--.— .--do. . . . . . .0076 4.24 6 --- D2 . mu L28 ~.6 176 .. —-. - --- - . .._do..-._ ------- 424 -.--.-- -— .— a5 .#7 5 274 .25 –: ~ ~::.--- .------ .--do . . . . . 4.14 .mw ----- D2 8.8 L92 -& o .024 10 .–.do. ----- ------ . . . . . &4 L u 6.0 .026 159 ------ --- -----— 10 4.28 .-:Y. .---— –: g ~:::-:.. ------- ----- D2 at! 202 –9. 1 .C&7 2s9 .25 16 .0101 .. ..do. . . . . . 8,89 .0105 .---.- ---- 8.4 .% &6 .OM 140 .- ..-. - .- ——- 15 .-—-- ...-do.-...- 420 --...— –. 16 ----.--- –— ,----- D2 8.6 2.11 -:$: 331 0 . ..-do____ .0162 6.80 .0132 ------ ---- 8.6 .84 :%! lm ------ .16 -------- ------- % ._Jlo ----- ------ 414 ------
I
----- D2 219 –12. 3 .086 84 0 .. ..do.-..-. . Oas -.10 -------- -.-–-– ,----- ----- k: .72 lL9 .694 --— E .–-do ------ .---— ---.---- --.---— .------- ,----- D2 8.2 2!29 –M. 6 .m 827 .—-. 45 .--do . . . ..- .----— –. 25 ----..-- —-- 8.6a . . ..-.. - ------ ---— 8.8 .m lito .111 Io2 .4 46 .--do . . . . . .Oxl -.--.-– ------- ------- 4.70 .072 .----- D2 8.8 2.25 -17.0 .0s2 327 ----- .----— -.26 ------- --—-– .-.do . . . ..- K72 -..:i T .--. . . ----- &4 .71 17.I ..- .-. 101 .4 % .109 -------- ..–--– -------- ...-do. ----- -------
I
.---— D2 8.4 229 –19. o .0f6 241 ----- 75 ------- –. 83 -------- -— ._-do . . . . . . 2.67 ..-:%- ------ ----- 8.6 .70 I&o .—. lW .2 75 .162 ------- ------- ---—-. __de____ .--. .— ----- D2 a4 –27 .loa WA .3 . Olm -.04 –-— ------ 3pI[t. .__.. L 41 .0103 ------ D2 M -40 .lm 2S3 .6 . owl -. m -------- -.–— .._do...-.- 4E4 . ON -. m --.--— ------- ------ D2 M 209 –5. 4 .104 .6 .ctzs .-.do.....- 441 .--. . . D2 8.4 2!M –5. 6 . 1G3 z ----- *.042 -.13 -.—–- —- .-_do._..- &51 ---:Y_ ----- D2 8.4 a.u .036 m .— 0.072 –. 18 -------- ------ ._-do ----- 400 ------- ------ D2 2.EJI -3 i ----- *. S2 –. 22 ------- ------- -..do- . . . . 418 ..-.— .-—-- D2 H Z64 -146 :%! E .--— ! 16 –. B –-—-- .----— .--do ------ 404 .--.--- ..—-- D2 8.1 264 –1s 6 .cw 382 ------ ! m -.27 ------- -------- .._do ---- ?bSa —-.
------ D2 8.4 2m –16. 7 .062 257 ---- (22 –. 2) ------- —-- .-do- . . . . . L72 –.–-..
— 1When tbe afrfotffeJnverted,a ndnus defiectkm ofthe flapindkatee that the fhp is detlecteddownward.
i ~ O(chordoft~ ofrfonwith fiP neutral- A refers to a chord defb?d ess he Jokdng tire Wremkks of the mean Me.
$Type of pressuredlstrfbntion. See reference8.
t TyL@ of scaleeffect on maximum Uft. M reference 4.
JType of Jfft.cuma@r as shown in the sketches.
A’fi /$ c ~Tnrbnlence factor fs 2,64.
7 These data Iraye ken ermceted for tlp et7ect.
~Angfe of zero lfft obtafned from Uneerlfft eurre appro.dmatfng ~fmental Ifft cure.
~810w obtabicd from llnear lift curve approxfrnatlng expei5mentel t curve.
lay oukfderangeofllft metWlent9 emeredfn thaw te3ts. Valueofci. gimn applle3 apPr0xirnete15 over entke usefulraw of IIft coefEeients.
n% -- m m-i.
11~ k tien abeut tfm aerodynamic center of the efrfofl wftbaut the flap and fs the average velne.
=ti...h u Velw of U, -{~ used fn eompntfng this ratii are tekerd mm teXs of ths Plrdn airfoil.
209142-4&26