Experimental and calculated characteristics of three wings of NACA 64-210 and 65-210 airfoil sections with and without 2 degree washout

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NATIONAL ADVISORY COMMITTEE

FOR AERONAUTICS

TECHNICAL NOTE No. 1422 * ~ ~EC 1947 EXPERIMENTAL AND CALCULATED CHARACTERISTICS @ THREE WINGf OF NACA 64-210 AND 65-210 AIRFOIL SECTIONS WITH AND WITHOUT 2° WASHOUT By James C. Sivells Langley Memorial Aeronautical Laboratory Langley Held, Va.

FOR REFERENCE

Washington August 1947 IJENARY’ cop’.

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LIBRARY NASA HAMPTON, VIRGINIA & P A TT AR!?

1A n UAL --m L,ANGUYMEMOXIALAEXONAWCW ~RATORY ..

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.:.;.,:; # EXPERi3@lWAL h) CALOUIhD CHUWTERISTWS (IF Tli@ii :iiiiiiGBm ‘: ““”k:’ “’:,: ‘ -, ,“ ..OF NM!A .6&210 AND ,,65,+10 AIRTO.%” dCTIh6’ ~.* , ; : ~ ~.

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An invest&ztiqn W6 been ccmdurkd. to djetemin6 sgme” of’.j &e” ‘; . .~; effects of airfoil section andWashdtit on We’experimental, &nd . . ,, calculated characteristics cf 10-jjerc&tWxLck wI-. ... @n&s:: of aspectratio’ g andI ratioof root chordto tip qho~:2 .,5 fiem,. tested .“ ‘ One @@ had WA 64-2J.o’ secthm and 2° wa@ou%$ the secpnd M NAOA “65-21O section@ andq? washout, @d ‘the W@* had: NAGA..65.-2U: ‘, : . “ sections” knd 06 washout. ~It was foundthatthe“’experimental c~ac$eristicsof thewingscouldbe satisfaotortly” predicted fro~ calculations based upon tm-dimensional. datawhen ‘l@ airfoil contours of theWags confomed to the tr@”airfoil “sections @th the same — highdegreeof accuracy as the twO-amO~iOti tidelii. Small constrwkbn emors werefoundto causelargedlscrepamcies in the .

valuesof max3qm liftcoefficient. .’ TM ‘most sigKL~iCant Off’OCt of .. ; .

changing the airfoil section fromm NW 61XZL0 S?ctfOn’to ~, k NACA 64-21o section was to increase the znaxiqm Mft “coefficient by .

about1.0 percent, although the abruptness of ‘&e stallwhs also ,.

increased.ne tia.it~oll Of 20 washoutto theNAgA65-21o. WIIM ,..

Increased the angleof. attack for zeroIlftas e~ected but was not mfflcient to improve material~v the sl+ing charactert.sties l .

The othercharaoteristtcs of the wingswero6s&rW.Uy. ui!fected by the changein airfoil section qr. ~y the addition of wasiiout.

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Ons meansof increasing the crltlcal speedof an kirplane wing F is to decrease the thichess of’ the airfoil sections.Decreasing , .4 the airfoil thiclmess belowaboutM! percent, however, reduces the.

maximum liftcoefficient of the section (reference 1). Furthermore, .

theuse of thinner airfoil sections increases the structural problems :: h encountered in the alrplans desi.g. As is usually the case, tn airplane desi~, somecompromises mustbe madebetween theseconf~lctlng . . . . . . . . . . . . . .

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L ikwderatlone.lb’ wYmlJlpl.e . an airfoil thicmsm represents a reasonable &m&romisebetween the ~rodynam~cand structural. considerations jmvolvea, In thedesignof a long-ran go, -.

hi~-s~ed atrplane.

Although the two-d,imsILsioI@, mmtiristics of sevciral thin dl?foil saotions havebeenpresented in ref’orenoe 1, verylittle da~ at relatively high Reynoldsnmbers have beenavailable on the three -dimensional che,r~tirlsti~s of wings inoorporati~ suoh sections. An Lmestigation-S *refare been conducted in the MJl&LOylg-footpressuretunnelto determinesome of the effectsd airfoil secticm and washout on the maxink liftand sWLlinG charactxmistjms of 20-~ercent -thick wings, Threewingswereinvesti - *W.: The firsthavin~NACA 64-XLO sections and 2° washout, the second hmi~ IWCA 65-210 xctiom m ao WWJHt# and the third, having MCA 65-210 secmicm and 00 washout. me phn fom of all threewi@s was typzcal of that~orwingsof a lo~-range, hi@- speed. airplaneIn that the a@ect ratio. was 9 and tie ratio of rootchordto ‘tip chordwas 2.~.

Presented hereinare theexperi- mentalaerodynamic characteristics of the threewin~s,toC@herwith theircharacteristics calctitedfm ~0-dwnsiow da~ accor~~r to the method of reference 2.

Cq!!l!mcmrms AND S’!lmom The. coefficients and synibols usedhereinare defined as follows(ccmsi~tent units): liftcoefficient % maximum valueof , C%kax CD ‘ dragcoefficient (D/@) profile-drag cosff’icient (Do/@) . .

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minimum valueof profile-drag coefficient C%ilin .’ pitohiqpmaentcoefficient (M’ /@6) h.

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lift L .; D“ ‘“ drag 3UCA~ NO, 1422 : (L/D)- vaheof ’ ratioof lift to drag Incwnml hag profile j@tching momentabout0.256 &p (9) @namic pressure of freO. .streem / ., win~ -a (24.94 sqft] - aerodynamic cliord ~.~:’+ (herein,6 = 1.769 ft) mean “.

mass’ density of air ‘ , ., .

., airspeed ‘ .

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wingspan (15,ft).. . . ‘ spanwise coordinate ,.

corrected angleof attackof root chord, “degyees an@e of attack, f or zerolift,de~eea, a(L=O) # R Re~qolds number (PV6/p) ‘ .“ .

M Machnwnber (V/a) ‘ coefficient of viscosity IJ SonicVulocity a ..

slope of liftcurveIn linear range, per d8gree .

pitchingymxnent,co efficient for zerolWt ‘(L=O) ., .

ti~/dCL slopeof’ pltchtng-uoment curvein lii’my r& f section-lift coefficient” .

.’ & maximumvalue of sectionlift coefficient ., . .

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The thrm wi~s were constmzcted of soMd steel and ‘were’” @metrically similar except for airfoil section andwashout.Ons windhad HACA 6k-210 secti.ans and i?o washout: the second had MCA 65-21o sections and 2° washoutj and the third had NACA65-210 sections and 0° washout.,The ratioor rootGhordto tiy chordwas 2.5 and the aspect ratiowas jl.- The sweepand dihedral a% the 0.25chord linewere0° and 3°) respectively. The winfjs with washout had unlfom twist aboutthe 0.25 chordMne inasmuch as COX’lW~OIldiI!,& elemsnts of the rootand tl.g sections wereconnected with strau’ht line s.

The wingsweremooth and’fair and cor@ormed to the true contour to within0.003inchoverthe forward 30 percent of thewlnG andwithin0.008inchovertherearward portions,The general dimensionof thewingsare givenIn figme 1.

The tests were conducted withthe air in the tunnel compressed to approxlmatily 34 pounds per square inch absoluto p-mum. The tests were made at a *nswk pressure of apymximately 05 pounds per square foot ) cor.rwpcnding to a Reynol.d.snwriberoi’ approximately A,hOOiOOO and a Mach nmber of about 0.17.

The aerodynamic forces snd xoments were masured by a shmzl- taneously recordin~; six-ccauponent baknce qystem. The profile dr~ was determined from the force test data and also ‘by thawake- momentun methodfrcmsurvuys of the air flowin thewakeof thewing at 19 spenwise stathns. The stalling characteristic weredetermined framob~ervations of thebehavior of tuftsattached to the upper surface of thewln~ hehlndthe 0.30chordU.ne.

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RESUUJ!S AND DISCUSSION All datahavebeenreduced to standard nondimensional coeff iciarcts.

Corrections havebeenapplied to theforceandmomentdatato aocount for the tareamd interference effects of’ theu&W sup~xxrt system.

Stream+u@.e and Jet-boundary corrections havebeenapplied to the, ‘ angleof attackand to thedra~coefficients, CkmqmWsonof’ Exger~ntal. and Calculate& Characteristics ‘1 Forceand momentcharacteristics. - The exyrlmental and a calculated lift,drag,and pitching-mment cham3ctoristics are shown In f~gures 2 to 4S k swmary of theda&ta in thesefi~s is gtvenin table1. The calculated chsmcteristics were obtained . .. ..

NIK?A TN MOc 1422 5“ by theuse of theme~od of reference 2, whichtakes into account the nonlinearity of the section liftcurves.All section dataused In the calculations wereobtained fra reference 1. In general, the agreementhetween theexperimental. and the Gdctitea character- isticsis considered to be verysatisfacto-ry~ This:~ood ameenent was possible, however, onlyafterextreme cue W- W- ~ ~~~ the’wing contours conformtothetrueairfoil ccmtours withthe SSJES highdegreeof accuracy asthe two-dhensional models. Thatthis extreme carewas necessary was indicated by ~reliminary testsof the NACA 65-210 win& in whichlargediscrepancies $nmax.@umlift coefficient werefoundto be due to small errors in construction} particularly aroundthe leading edgeof thewings.

: The mahdlscrepancfes between the experimental. and cal-culatid ch&acterlstics occurat theupperend of the law-drag ren~ where the calculated curvesshowa moreextend.ve low-drag rangethantie probably”due to the spanwise experimental curves, This.effectis spread of tra,sltion.which was not takenintoaccoynt in the calculations.

The reasonekd.eness of $his’explanation is Jndicatedhythefactthat the low-tiag rangeobtained fromthewake surveys a@?eesj in 6SnWal, withthat.ohtai.md fromthe forcetests. (TheproiUe-drag,coefficients obtained fromforcetestsweredetermined. by subtracting the calculated induced-drag coefficients fromthetotal-drag coefficients~ The difference in the extentof the low-drag ran~ Is alsorefl.ectsd in the vsluesof (L/D)~ sincethesevalues, in ew?y case,were obtained at the upperend of the low-drag r-. This~sult emphasizes the needof preserving laminar flqwas far as possible inoriler to obtain hi@ valuesof (L/D)~.

, i .Stallind characteristics.- The stalling characteristics of the threewingsare shownin figure5. The valuesof liftcoefficient .

shown-~reobtained with tuftsin placeon thewing. In orderto “, predict the stdl.ling characteristics of the Wiwsj the characteristics calculated according to themethodof’ reference 2 and presented in fi~e 6maybe used. Thisfib@reshowsthe a~emwise variation of themaximum liftcoefficient whicheach section is capable of.

reaching in two-Umensional flowaad the variation of. section l$ft coefficient for thewingwhen somesection firstrea&eq itsMSXUUUM value, According to reference 3the maximum lift coefficient of the wingis reached when tho curves first’become @.n&nt; thepoint of tangency of the two curvesindicates the symrise position of We initial stall,and the rateof divergence between the ctc?ves serves as an indication of themannerIn whichthe stalls~eads.

Frornaccmqyxrison of figwes 5 and 6, it canbe seenthatthese wingsstallapproximately as predicted..

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. ,. Ei’fecrt of Airfoil Swtbn ,, A comparison of ftguree 2 and 3 shows that the minimun I pzktile -drag mefficien~ of the NAOA 64-21o WIW iS d.I@Itly hi@er (about C) l 000k) tha that of the I’?MA65-22.0 wing and tho maximum .lii%-&ag ratio is corre6pondi* lowerthanthatof the NACA65-210 winG. Themost E@xL?icant effectof thedifference in airfoil section is,howevwrJ the approximately U1-percent increa~o in meximm.m liftcoefficient for theNACA64-210 wingovwrthatfor theNAM 67-21o WiIU m FZWM figure 5 it may he seenthatthe stallof theNACA 64-g10 wing begansli@rliQ farther idm,rd than that of the NILCA 6J-21O wing but was moreabruptandwas accompanied by a @eater 10SH in lift.

Figwe s hdicates that the” stall of the NACA64-210 wing should . . ‘have be~unsli@M.yfazzther outboard tha?,?, that of’ thO lfAOA $-21O whg but theU.fferences in speawise position in eithercaseare indicate that beycmd uaximr: lift there small . Bothfigures is sli@tly less stalled area at the wing tips of the l?ACll 64-~l@ wing.

Effectof Waslhout ,,.

Exceptfor the expected change in the+- of attack for zero lift,thereareno -practical. differences in the characteristics of’ the twolWiCA 65-210 wingswithandwithout 20 washout(fi&s, 3 .

~~ 1~) , High-speed testsreads at theAm9sAeronautical Laboratory b of shntlar wfngsalsoshowed negligible effect of 2° washout,A larger, cmtmnt of washout would, however, probably make somedifi’er- ence,but the amountof washout #at COUM.he’ tolercdmd without ., introduc~njq harmhil. effects at h$@ Mach numbers is not knom.

The 2° washout was not enough. to imprc.ve materially the StSJL@ c~acteristics of the NAM 65-210 wings, &Lthough the .

Spanwise position of the incipient Htall was mowd s-what inboard because of thewashout.A;larger. mount of wa~hout should etfect a signlf icantimprovmmnt by movingthe stall farther iritxmrd.

, .: (101’V2LUSiOILS From the’ r&lk~ of an Inmpkiga.tion of thee~rimental and calculated characteristics of thrq.e l@percedMlxl.ck tapered win@, thefollowing conclusions may be drawn: 1. The experimental charaoterietics of we wings cOtia be # ‘satisfactorily predicted fromcalculations basedupontwo+mensional data when the- s&foil contours of thewi&s confor&dto tk true -— airfoil contours with the samehighdegreeof accuracy as the two-dimensional mode:ls from whtchthe data ,were obtained. Small errors in construction were found to cause *ge discrepancies in the values of maximum lift coefficient 2. The‘most slgnlfic&rk effectof changing the airfoil ‘‘ section froman NACA65-210secti~n to an NAOA64-2x0section was to Increase them+clmum liftcoefficient by about 10 percent, although the abruptness of’ the stallwas alsoIncreased.

3. The addition of 20 washo~t to the NAOA 65-210 wing increased the angle of attaclc for zero lift as expected tiut was not sufficient to Improve materially the stalling characteristics.

4. The other chract.mistics of the wings were virtually “ =fectetl by the:change in airfoil section or hy the small amount “ of washout.

_ey Memorial Aeronautical Laboratorsr — I!ktional Ad.vlsory Committee for A&onauttcs lku@.ey Field, Vs., Auguste, 1947 .

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1 Abbott, Ira H,, von Doekoff, Jllbert E., an&Stti~rs, Louis S,, Jr.: Sumaxy of Airfoil Data.

~QIMR No. L5CC)5,1945.

2. Stvells, Jams C., andl%e&, Robert”E.: Md2md...fcr Calculating .

Wirg-Chazacte~istics by-id.fting-Line Theory ti8i.ng NonUnm3r- Section Lif*Dsha.

It4CA ~ NO. k26~, @K( s 3. Anderson, RaymondF.: .Detexqinathn ofthc Characteristics of Taperd WirM;S. l’W.CARep.llc. 572,1936.

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Calsumtd WJ.o@ataa Oalclaataa 1.36 1.35. 0.0%” 0.083 38.8 38.8 -1.0 -1.1 1.21. .085 “1.0 41.0 1.24 -1.0 39J+ . .0e5 1.23 .085 -1.5 38.0 41.6. l o& 1.22 -1.3

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H* Oalcuktaa OalculAtea 0.0045 0.0043 0.0044 -0.037 -0.040 0 -0.010 .0041 .0040 -.035 .0039 -.035 0 -.OI.O .0041 .0044 -.037 -.035 0.004. -.006 R4!cImumvmmr om4ETlmm AEmfmms 11.43 —— ..- - T :, 180.00 i I I I ——_ —- .

‘—. .—— I NATIONAL ADVISORY COMMITTEEFORAERONAUTICS . .

Figure 1.- General dimensions of 10-percent-thick wings tested in Langley 19-foot pressure tunnel.

Aspect ratio, 9; ratio of root chord to tip chord, 2.5. (All dimensions are in inches. ) l g .

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/.4 /.2 .8 .6 CL .4 .2 :2 -./ 0. .(w4 m 0/2 002#z% #./O-4 048/20 GDO C’n a Cm Figure 2.- Experimental and calculated characteristics of”wing having NACA 64-210 airfoil sections.

M x 0.17.

Washout, 2°; aspect ratio, 9; ratio of root chord to tip chord, 2.5; R % 4.4 x 106 ; .

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/.2 /.0 --- Csleulatad - +’ I 1%1 .8 .6 I -o i Inm2e tams cl .4 .2 0 .004 008 .0/2 0.02 #DfS,@.iO-4 048/20 7/ co CD # cm o .

Figuxe 3.- Experimental and calculated characteristics of wing having NACA 65-210 airfoil sections.

Washout, 2°; aspect ratio, 9; ratio of root chord to tip chord, 2.5; R x 4.4 x 106; M z 0.17.

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,, -,,,,, ,,, #m -.;* /.4 /.2 /.0 , , , # 1 1 “ 1 / ‘o t .8 I I I # I i a I .6 q -O Fbrcet4 A .4 .2 0 Do4a08t2’2 7/ 0 .02040608./0-4 048/20 u- CDO c~ c~ Figure 4.- Experiment&1 and calculated characteristics of wing having NACA 65-210 airfoil sections.

Washout, 0°; aspect ratio, 9; ratio of root chord to tip chord, 2.5; R % 4.4 x 106; M z 0.17.

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Rough flow Cotnplde doll ‘d /

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Infermiffd still

Cress flow

a:t3. o” +-1.18 a.~& ~= .99 a= 11.9” CL=1.10

a= 135”

~= L21 a=/5.0* cL,/.3~ a= L23” CL= [06 a= 13.9” “ c’= Lii?

ad5.0” c~:h97 a=l.i4” CL=100 a= 14.3.

CL= ,98 (h) NACA 64-210 (cJ NAC$~Y%J;( WM79; 2*washouf.

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Figure 5.- stalling W.racteristics of wings of aspect ratio 9 and ratio .

of root chord to tip chord 2.5.

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NATIONAL ADVISORY ~ COMMITTEE FOR AERONAUTICS . “8

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(a) NACA 64-210 sections, 2° washout.

Mt coefficient and section lift coefficient at maximum - &

?igure 6.- Spanwise variation of maximum section wing lift coefficient forwingsof aspect ratio 9 and ratio of root chord to tip chord 2.5. R s4.4 x 106= .

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, (b) NACA65-210 sections, 2°washout.

Figure 6.- Continued.

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