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FOR AERONAUTICS
TECHNICAL NOTE
No. 1151
OF DRAG CHARACTERISTICS OF
PRACTICAL-CONSTRUCTION WING SECTIONS
By Joti H. Quinn3 Jr.
Langley Memorial Aeronautical LaboratoryLangley Field, Va.
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Washington
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NATIONU ADVISORY CGM&TT~ .-
‘IXCENICALNOTE NO.
SI.MMMIW03’DPAG CMRL2TEKC3!.TCS 03?
E’FUCTICAL-C!OEX?RUCI’1ONWING SECTIONS
Ey John H. Qui~, Jr. ..
smMf%RY
The effects of se~erd parameters on the dr~ characteristicsof practical-constructlctn~ sections have keen considered and~=w3~5a. The effactm cors:derwi were ‘dose of suz-facerou@nc3ss,surface wavinese, ccmyressive load, end de-icers. The data wereobtained trcm a number Gf tes’% +inthe Langley two-dimenstcnallow-turbulence tunnels.
——-.-. .
The sectim drag coef’fici~ts of -p~e,ctical-constructi.cnw-sin the “as-recei~ed” ccmd.~ton wero cn%6n aJ3high as 0 .“OO~OEt_ ‘- --
&Reynolds nrmikrs of 20 x l.O. W&n k@r jo@ts or s@facti tilfairwwoccvrred in a region of ncnmally Laminar fiow$ dec”rea~esin stictfcm-drag coefficient up to ~0 percent could b-eoltdti-e~%y a”c&6T.&tiofiof surface finishing and “?airi”ng.~“ sahe caiieejnd=l.y hal? thisimprovement waa due to bet’%r surface fdrn~si .“ The-&b*- Z3”fsmo6fi”
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w:ngs with thick skin having spars placeq at or kehtnd the mos_trtierwartlposition at wM.ch Uumlner flow m5@t be eqieti~d ~~oach~dthat of fair and smooth airfoils of correspmding se6%30ns. &m6 -queatitati.vedata were obtiincd ad. indicated the effects ~f “wa%%s’-in the laminar-flow r~@xl of E3moothpr~g-t=ii~-construction W*.S onthe ReymGlds nuniberat which yrtimaturet.rmmi-~-ionwouti ‘oe~-ur~”‘“FG%’–Rejnolds nudmrs up ta x x 3.0=,a few cxemple~-ax%-gfven of inzrf%cewaves on N.ACA6-series -foil gect&s that did”no% c“au%epr~~tie
.
tram iticn.. .
As a result of th~ constructia irregukrtties existi~m onwiqy as received from the msx.mfacture~,the dtffercnces in &a&
.----—.
uEually associated with a:rfoils cf d:ffaent se-ries‘imre”not ohtalmd.Ccmbinatlms of glaz+ing,painttng, or minor refairing of the surfaces,however, w6ra suffictent to prcch.zcesecticn ?ma,gcoeffici~ -
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a.pproaohimzihoee for tair and smooth airfoils of correspond~u sections .at Rqnolde nmibere up to approximately 20 x 106.
Loading a wi~ in comprassloa until some slight perzcsmenteotof the skin or rivets occurred [email protected] or no adverse effect on the
I?JKMTN No. 11512
dragtruewingdrag
characteristics of two wing sections designed W retain thel.rcontours under loads usually encountered in flight. While thewas under load sufficient to produce.such deformation, however,coefficients ae high ?s O .O@O wero obtained at G Reynolds _.
number or approximately 24 x 10” aB compared with a value of 0.0045for tho unloaded wing at the sme Reynolds nunibor.
Airfoil secticms having thickness ratios of approximately15 percent and equipped with leading-edge de-icer baota were frondto have section drag coefficients of approximately 0.0070 at Reynoldsnumbers between 3-0x 106 end 32 x 1~ . This value of’the sectiondrag ooefficlent appeared to be independent of’the airfoil sectionupon which the de-icer was mounted.
IIVIRODUCTTOIV
Numerous investigations of airfoilPractical-coneim.mtionmetiods have been
sc3ctlons-builtby Variouemade in the Langley-two-
dimensional low-turbulence tunn~16 to dotmmlne the effects ofconstruction irreduhrities on the aero*mmic characterif3ticsofthe airfoil sectdcms that each n~del--representid. The results of thotests were useful in estimating perfomnanoe characteristics of theairplane for which each installation was being considered, but noattempt was made to correlate the aerodynamic characteristic of thewing sections with the tyye of constructicm employed.
In the present paper the data obtained from the tests have beencollected and analyzed to find the effects of several parsmetere onthe drag characteristics of practioal-constructionwings. The effectsof surface roughness, surface wavinese, compressive loa~, and de-icerswere considered. The drag characteristics of the models, which repre-sented both NACA 6- and 230-series airfoil eections, were obtainedfor various surface conditions. These surface conditions generallyincluded the original condition as received from the manufacturer anda number of Improved conditions obtained by glazing, sanding, painting,or by a combination of these processes. Surface-wavinessmoasurembntswere made more recently on several models and the drag and wavinessmeasurements ware correlated wherever yoseible.
c
d
SYMBOLS .-
alrfoil chord, feet
difference between gage rea~lng on airfoil surface and ona flat plate, feet
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“
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●
r
NAC!AT!?Ho. 1153 3
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‘.
d/c
s
cd
GI
Czi
R
%
x
6
‘5
u
U.
s
E.0
P
!10
wavine6B i.1’ikx
s~cbicn drag ,cceffl.cient
section lift coefficient
design section lift coefficient
X@nolds numkr bc.sodon wing chord
accf31erationof ~avity, feet Qer second,pezzsecond
distencG alo~ chord from leadiag edge, feet
effective thlckmss of bourilarylayer; thiolcw3ssto pointwhere velocity inside bomdary layer is equal to 0.707 ofvelocity outsidu boundary Iayor, feet
Reynolds number based on .d?foctivoboundery-layer thickness
IOOEL1 pelwi~ outside boundary layer, feet per ee?~d
f!reo-strew,mvelocity, feet per second
() \Ho-ypressure cooffictont —-
~Q
freo-strecqntotal pressure . —.
local static prtiesuru -. -._
frae-stream dynamic prccsure
McmdIs —
‘I& modeis tested were built by practical-construction methodsand were of ~-foot span and fnn 6- b 8.~3-foot ~holyj. Cficrdwlsostifftiners,spanwlBe stiffeners, or c.anbinationsof th9 two wereused, and the models were of the Gingle-; double-, or +triple-@rt~e. Both NACA 230- and 6-series airfoil sections wera rqrosented. .Explamatione of the aixxfoildes&nations are included in reference 1.
NACA TN No. IJ.51
The ort@iial condition of tilewing es received from themanufacturer and also the vszlovs inq?rovcdcondition are describedfor each Ddel where data for the mri.oue surface finishes aropresented. These improvod.surface ccmditions were o%td.ned by meor ?mwe of the fcill.owi~finiehind procedures:
Camouflage painted: 1%.$ntedwith sp.tietic-enamel ca.nmu-fl.agepaint giving a m.mface condition similar to that obtained byprocedure 5 of reference 2.
sanded: Suzfwo awil,edsufficiently to remove paint specksand other similar excrescencor3.
Glazed: Local defecha such a~ nioks, dim@es around.rivotsj end seems, fil.lodwith ypoxylin putty end sanded smooth.
Ps,intod: Painted with gray pz@sr surfacer and sandedsmooth with No. 320 Cd301?l.UYhWl@@cl’.
I’aired: Modification to surfaco either by extensiveepplicaticm of pyroxylin putty or robufldind to reduco MO numberand size of larger surface irrcgularftdea.
In tho present papa’ tho term “roughness*’is used to donotothe pro6ence of local nicks or Qcratches, opon seams due to chord-WISO or spanwise joints, dimp].esaround rtvds or.screws, paintspecks, or other similar projections. Tim.Mm “waviness” is limitedto Uhose wrinkles in the slclnthat preoent-gcmtl.edeviati.~s from afair surface. A surfuc~ ie considermi to be aeK@_cd.ly fair andemooth when further decreases in the amount of surface rmgiumm andwavizlesaproduce no chsnge in the aerodynamic characteristics.
Descriptions of tie @ale, a list of the surface conM.tinnestudied, and an index to figures in which data for &o varioussurface conditions are contained are presontxxli.ntable I fnr themcikls considered herein. “
TEST METHXL03
The tests of the Tractical-constructionwing mcihle were tiein the I.a..r@eytwo-dimensional low-turhu.lencotunnel (desi@ated L~)and in the Lan@ey Lwo-dimensirmallow-turbuloncoproseuro tunnel.(designated TDT). These tunnels have test soction6 3 feet wide
by 7; feet high and were deai@ned to teet @ols com@ete~ ~pcmning
the jet In two-dimensionalflow. The ln?.rbuloncelevel of these
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tunnels amunts to on3y a i!owh.mdredti.~of 1 ~ercent md is_—— .-.considerably below t.c.tat which my effect is apparent.cm the”-critical Re~olds nuriberof a sphere. l’estsinthe~~fbe *“under pressl~ws ran&ng fi’OIIL 14.7 to lx ‘pO~n@ psr.sg~e .&@ ._alwolute; theref01%,t~ increzmi~” ‘&e tunnel @f3su&i ‘lI@:.I&yn.otinumbers nay be obtained et rek ktivelylaw Mach nmibers. The khchnumber of the tests was in no case greater tlmn 0.2. In them .—
tunnels, lift is measumd by “integratingthe preosu~Os along thefloor and coilinG of the tmrmel test section C@ dr~ is mo~ured. bythe ‘wake-m.zrvoymethod. The dzze~coeffic@@p- artitiudly obtained ~at a spauwiae ~osition selectciias a ieyrescntitiv”esectiti of ~qi _wfn~ from a n~ber of spmwl.se surve~s at a “l&wlift c60fficient.Moro detall&i descriptions of tho methods tied in obtaining andreducing data in time tunnels tie containod in r“eference 1. . . ..—
Suzrfaco-wavinessmemmxrmxmts for the ,wfni-tunnelmodels wereobte.lnedwith a sta.nh.rd.Ames dial ga&e mounted on legs syaced
21: fnch~s. ?&e readings mm rmlu.cedto &lmensicmb3s form by-.---.+--_-,.... .. . ____
subtracting the re- of tie gage when placed on a ‘flatsurfacefrom tlaereadings obtainod with the fiage‘iInvqious ~os+ti~- alongtineairfoil surface and divi~ the dlffbroncobY”&a@ airfoil chbfi.—. ---—
RESULTS AND MSCUSSION —.
In the endysis of the mww of aurfeco roughnem andwaviness, the surfaces wera amumed ta be so”tiooth that’the -‘differences observed between tie ?mas~”~ bags and the dra@ offair end smooth mm?els were related directly +~ me re~=t~ve extin~of the laminar ~nd t~-btiL~t bounm l~+~i-s. ‘2he effects-of ~faceroup$ness or wavin+88 on draG therefore can be interprOte~ essentiallyas tileeffect of this i*0@ile6f3or wu.v~mtisti& the ptisititiof thetransition from the laminar to the turbuient‘@er.- —
h order to derive an spproxhmte relaticm between * secticndr~ coefficient m.d the position of tr9nsiticm, section drag c&Zt-ficients have teen calculated by the metkod of reference 3 for theNACA 66(=5) -116 airfoil section at a section lift coefficient of 0.1
and a Reynolds nunilmrof 20 X 10s for aesti positions of transitionrangi~ from O.lC to o.6C. (See fig. 1.) Those calculated valueshave been used Lhrou.ghouttileanalysis when ~ e~timte of tietransition point on NACA 6-series airfoils wkm required, sinco theVeriation shown in figure 1 is “Kmught to @ reasonably repre-sentative of the airfoil sectims for which data ore -pros&ted h~rein.The values of the cection drag coefficient-found ‘for trtis~m .“:‘—-
6 ~~:. ~ NO. ~~1
at O.Wc or O.60G are probably slightl~.h$~er ibn those of fairand smooth NACA 65- or ~-~eries airfoils, respectively, bocaumat Reynolds numbers up to”.qproximately 20 X loG trmsitlammu.ldprobably occur eliglhtlybehind tiie mfnhxn pressure point.
Effects of Surfaca Conditlone
Surface rough ess.- In the.cozmidoration of tie effects ofsurface rO1.@ll(3SSon tho dr~ chu’acteris~iCS Of puCtiCa&construction wings, the eepara~e cffects of various steps in “tiofinishing proces~ have been determined. Pt~otographsof models 1to 6, which are NACA 6-serim eiri’oiisections, [ne presen’todCMfigures 2 to 7. !lhedr.ngcharacteristics of thosevarious surf’acoccmd.itions am presontid.in fi~m?e
Frcnnfigure 8(a) at a I?aynoldsnumber.of 20 XfOllowi drag characteristicsmay be obtained for(NACA 6f216) -3( 16. a) (approx.} eirfcd.1Bectim):
fw?-y
2
3
moaols with8.
2.0G themodel 1
Surface ccaiditionI
Perceritqy3cdimprovement— .—
@w@inal.,camouflage pnintod;discontinuity at’frmtspar (0.12c)
Upjer surface glazed ommfront-Syar; lower surface@azed to front spar
Upper surface yaintod toO .71c; lower surface ‘painted to O.22c
Both suz!faces painted t-cO .71c
0.0086
.0070
.coy3
.0052
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19
33
40
An irregularity consisting of a rather lar~e flqt spot existed attho frent spar (2.12u) on both eurf.acesin ~ho [email protected] flak s,potwas detected by rockin~ a atiaig.!htedgeover thesurfacas in a chordwise direction. !I!hL.large reduction in dmgobtained from stop 2 was prcbably duo to a partial ftiirtilgof th~flat spot on the upper sqrface. Transitdcn moved ihwnstrwm butstill occurred forward of V.m minimuu pr.ossurepdint as a resultof the flat spot. Local glazing (step 2) }@ palnti~ the modelsurfaces (steps 3 and h) are not thought to il.tertho surfwm v’
wavinesu a.p~reciablybut rather to elim.inatalocal nicks, dimplee,
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W31 TN w. 1151 7-
Seamc, end &cratdi5Ei. The final val’leof the section ** coef-ficieni of O.(.X3-2Cb+&ned with stip 4 corresponds ta %m-itim ...at epprcixi.matelyG.43c, or O.OTC ahead C5 the dee@n pasit.lonofminimum presmre on an NACA 65-series airfoll mcti.on. Stice themodel surfaces afti~rstep h were,mooth tindthtimiddlo spar wuilocsted at O.45c. the rem’ nirg unfairness near the nose of thu
nd.d appaarud ‘A be responsible for tie Pr-ture ..$rq~:tign.__ _ . ..
The folkwing t@le dmwe the inprova.wnts made on model 2(MICA 66(235)+!14(CLppKCC.) airfoii eectlon) at a Reynolds number Of_20 x 106, as obtained z~om figure 8(b):
-. ——.
step Surface conditicn Pcn-centagecdimprovement
1 OriSinal,unpainted o.Oq’o -----------2 Glazed snd pzhlt@ .oo~~ ‘ 213 Refalxed ‘ .0035 50 1
Tliedxag was redxced ~0 percent, althongh a reducttan of @21 percent was obtainod by emootllingKte suxfaces. In tho unp~intedcondition, the 89cti.ondreg coofficfont Gf O.Oon corre6pcmdB totramition at approximately O.24c. Figure 3 shows that nmeroutidtipl~~ caused.by the rivets e.tistad in iih~sl~n. I!heBetimple~were probably responsible for trnneitlmn appro~tely O.10C aheadof’the flw.ntspar. Glazing and painting the tiol reduced.t@sectim dr~ co.ef?icientto O .0055, or naved tr.wsikim to apprti- -mately O.@c. T&mmition at ‘&i= point was prob~~hlydue to unfair-ness at the front spar. Refairing the model evi~emtly removed tieirzqulal’ity at the front qar an? the =ection drag coeffi.ciantwas reduced to the value of O .0035,or &pm’oxtitely tl:esam asthat of a fair end smooth model of the s~” section.
The drag character~stics of model 3 (NACA 66(217)-u6 aizyoflsection) ~e presented In figure 8{ c ) for e rengc of Reynolds
unltiers and in the tol.1.owi~table for a Reynolds numlmr of 20 x U :
P-’ I Surftit%~onditionI cd
12
3
i-
Ori.ggnal(bare-@tal skin) 0.006!2Glazed to spar joint at
OC3C ; .0053Glazed and painted over
spar joint .@&F~tire surface painted .0042
.W.rtzy refaired .qo401Percentageimprovement.——. —,—— .------------
n
2932
6
8 McA TN No. 1151
The Dectlon drag coefficfentif the model,in the original (b8ro-metal) condition, 0.0052, comwsponds to @ansition at appro.xiwtd.y
-.
o.32c. ~PbS and 10Cd defects fcnward c~ftoe #p~~ (fig. ~!).
prcbably caused transitlm at that point, Tho @.azing of thesurfaces foz?wmd of tha spe.r(step 2) rcduceflthe @a~ 1.1~ercsnt;tho secticm dr~ coefficiat of Cl.00~’jt-oaespo~”~ transition -
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at about O.@c. Glazing .md.ylnt,ing ove~ tiLeEpar joint (step ~)dticroasodthe aectim dwq coefficient to O .00&k, or moved tramittcmto approximately O.50c. i%inti~ lticontiro model.mrfaces (step 4)brought :!’boutl.ittlofurther tmprovmaent. Sariswavineas at the sparjoint at 0.32c (tablo I) wus probnbly resgon~iblc for yrl=turotreasition on model :. lib fina,Lsoc~lon draflCocfficiont ot 0.00@,however, shows t%at the we.vlnossdid not oauue prume.turetmnsi tion W’up to appr@mat31y O.5ZC.
[ (!)ke dreg characteristics of moiicl4 MM 66(215) -u.6
)
= ~,Oj Cz~ = 0-2‘1afyfri~SOCti1013(; = o.(5,Cli = -o.I_J -
arc prommtcd IKIfigure 8(&)
and in the following table at a Reynold= number of 20 X 10=:
7-=== C-tion I CL% J&%k%——1 Ortginal - painted wi’dh o *CX355 ------------
zinc-chromate primer
2 Painted .0040 m
3 G.~~e& ,r-J@~ eL *
A total rduction in cection dreg coofficfont frcm O.00j6 to O .O@10,or ZYJpercent, was obtai~iedby smoothtiu tilemodel surfacoa. ‘Ihemxlden increase In section drac ooofficisnt at a Re~olds numhw
of 13 X 10G was t&s climia~ted, F.Sshown j.nfiguuo 8(d].“ Ra~ldincl”0aE30sti SeCtiOn tia~ GOOffiCfl.GntWith Reynol& Illtiey,t3~l~to that shown, em um.mtllyussoclated with surface rwghneB.~. Localnicks or depressions nar the rivets proba~l.ycaused prmuaturetransition at a Reynolds nurriberof 13 X lo in &o ~mpainted contlltlonbut were not la.r~eenough to cau~e prcmatme ti”~DitiOn at lowerReynolds nimibers. The flush riveting m this modol was unnsucllysmooth. The final secticn dr~~ coefficient of O .OdtO is h&@er
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NACA TN No. 11~1
TJIanthat of a Yair ad smoothBecause the spar on this model
~waviness at the spar joint was
9
IfAcA66-series airfoil section.was located at O.@c (table I),not likol~ to be responsible for thio
discrepancy. Dev~atimm from true contour in both &o chordtise andspanwise directions, as shuwn in figlwe ~, the~fore, were probablyresponsible for the slightly hign drags in the finished.conditicm.
(
The section &r
T
coefficient of 0.0037 for model 52 = 1.0,
)0“21 airfoil sectionNACA 66(215)-u6 ~zi =
LafOuna= 0.6, Czi = -0.1~
at R = 20 x 10= (fig. 8(e)) is nearly tie ssmo as that of a fairand smooth 66-~eries section, and consequently little or no improve-ment was Wdo by painti~ and sz,nding. The spar location at O.&lc,combined with the uae of a thick skin (&bJ.o 1), prolmbly tiepossible the rcmlizaticm of luw-&ag characteristics to higherReynolds nunibezwthen he,vebeen fourd witiimost m@els having spars10CatCd fer’tiovform.rd.
Ve.rlatlonsof section dr~ coafficicnt with surfaco ccmditionfovmodel 6 (NACA 66(215)-11.6airfotl secticm) g.r~shown in thefollowing table at a [email protected] of 20 X 106, as obte.lnedfigure 8(f):
1
Rep Surfaco condition ~
1 Original.- coverod.with.fab-’ric surfacer
2 Fabric surfacer sanded
3 Snrfacor xmmoved
4 Glazoa UY tO o.15c
5 GlazrxiUp to 0.45c
‘=ES==I0“m661---------
1.0060 9
.0072 -9
.0072 -9
.0066 0
from
NO larce dec~eases in section @m-z coofficiont woro cbtainod byimproti-n~tic.swf:.ce finish of m-del 6. In the best conditi&,that is, with f~bric sur:ac~r samled, transition probably 09curr6dat Eip~rO~~Lh21y ().3’5C,or 0.2jc Amid of thc3de@gn positim cfminhum ,?re3blZa. me mrfac~ rmtmri:.1,which c~-istca of fabricdopd to +hs wtal skin, evidently mackcxlccms:derable unf=irnes-~for in ti~cbare--metalcontiiti.ontho drag wns 9 porccnt htgker than
10 NACA TN No ● 1151
that tor the model in the original condition. T!bf3drag co.sf-ficiont of 0.0072 for flteps3 end 4 would correspond to transitionat approximately 0.21c. Glazing to the rear 8par (step 5) 3W6ultedin a section dr~ coefficient that would correspond to transition atabout Ot28c. The model surfaces in this”oasf3were very #mooth; the
v extreme surface waviness-of model 6 therefore, was probably rcsponsi-blo for the high section dreg coefficient@.
The precmiin~ observations of the decrease in drag c~qcd. byimproving the surface finish end.fairnesp of practdcal-constrv,clxl.cnwin~u at a Reynolds num?mr of Xl X 10G we suxmnarizedin the followi~stutemmts: When spar joints or similar swrfaco irregularitiesoccurred in a [email protected] of nolmally Iamj.nurflow, the section drag coef-ficients of several NM2A 6-series airfoil sections M received fromthe manufacturer rauf;odfrom O.0062 to 0,0036. A combination ofimprovownt in surface ernoothnes!aand fa@ness obtained by glazing,ptintin~, or tinor ref~d.ringreduced Lhoqo section tiag coefficientsby an amount [email protected] from 0.0022 to 0.0035, depending upon the valueof the oritytnald-rags. Tests of twomod~~ having thick skti.sandqars placed at or btihindthe most i“()~:mti ptmitfon at which laninarflow might %e expected .yieldedsection ~ag coefficienL~ ve~ C1OBOto those of fair end smooth airfoils of correspcnd,j.ngsections.Elimination of’’tinorsuzzfacerou@mms by local glazing and.yatntinehelped to lmi.ntafnIihesovalues of tie section dr~ coefficient overa rather J_arSerange of Reynolds nuniber. Glaz@ and painting thesemodels did no’t~however, eliminate the adverae offucta of surfaceunfairncms ou waviness whore it existed, altinoughthe Etevorityoftheso effects was usually lessened.
.
—
.
.
@face waviness.- In the considorablon of the effects ofsurface wavine~s ur-the dra~ charwcterlstics of airfoil aectiam,the eff.ecti~of rou@mes6 have been oliminatm.1by using data forSmooth modols Only. ‘Thetypos of waviness invzmtiflatedwere thoseassociated with shoi’t-wav~-l~th wrinkltisin tho airfuil sMn undwith d.oviationsfrom true contouzzover a leqe part of the chord.The wrinkles, or waves, were dctoctod bypassing a surface gago overthe airfoil surface to obtain the wh~irl~g~ index d~c at a numberof chords?i.selocations. Any deviaticm fram a fair cwrva In the plotof wcviness Index ~ainst chordwiso yosition is en indlcatlm of asurface wave, albho”ug.lhthe w~vines~ index does not give directlyeither the length or ma~itude of the wayo. When the syacing of thelegs of the ~a~e is approamately a ccnstant frection of tho airfoilchord, howevor, the dovie,tim of the choddvise vsriakion of–thewaviness index from a fair curv~ iB a &ati.sf&ctorymans of comparingthe relative waviness cm di,fferontairfoil models. Deviaticms frcm
s
true airfoil contour ovor a Ie,rgepart of the airfoil chord wereinvestigated in one case by checking the modol contour with a tqmplet.
.
NACA TN HO. 1151 U
Feeler gages inserted between the templet and +fioairfoil surfacewere used.to measure the d,evie.tianfrom tho true contour1
The surface wavineas on two mcdels was reduced beyond tiepoint where an effeet on drag was noticeable. The two mciiclsweremcniel7 (tie NACA 66(215)-114 airfoil wc M_on) and model ~ (+&eNACA 66(2xL5)-116 airfoil sectilm), The drag characteristics ofmcilels7 and 8 could thpn be compared with those of otimr smoothmodels of similar airfoil ~ecticm to detarminewhether the dragcharacte~istlcs of the other nmdels were adverssly affected bysurface waviness and, if’so, to what extent.
A ploto@m~h of model ~ i.cpeaented as figure 9. The dr~characteyistics of this model wi+~ -bwoconditions of surface wavineEMare presontid iu fi~e 10, and the wavinoss measurements for thetwo surface conditicxm =W presented in figq.nwn’. Almost nodifferonco was found in the L-ag characteristics With the two
.,
waviness conditions, althou@ ins+mctlm of figure D shows that-iii “”tie faired condition tie model sw.’facesweco conoiderab~y moro fairthsn in the “as-received’icondition. Because a marked reduction intho surface waviness thus had no ap~~ont effect on the tio.gchar@cter-isticg of model 7, it W+M tJhou@t that transition probably mo~os ~ .-
foiw~’ as tie Reynolds number incras,sosoven if no waves exiet. Inordor ta invostiga+fithe possibility of this phenmona, dr~ co~f-ficionts wero ceJ.cul~tidfor sovoral F6ymolds nunibersby the methodof refcrenco 3. For them cal.culatlms it was assumbd that trangltiauw~la occur at a constent value of 11~ (Reynolds number baaed cm tb -effective boundary-layer thicl.mess) unless tie ~rticukr value
of Rb choson occurred behind the position of minimmn press~e.Estimation of the transition point in on adverse pressure gradiaut-israther involved md was not considarod of sufficient intmost. in thepresent paper to be included, llm positim of trsasitlcm was eab~tedfor several asmmed valueg of Rb botwoen’Cm arid 8500 by usc of thefollcrwi~ equati.onobtained from refwonce 4:
The use of a ccmstant value of R5 of 8000 was found to providethe boot over-all a~eement betwcmn tinecalculated and exywriMc.uMsection dzz~ coefficients. A.lthou@ the cdculnted-dr~ andoxpqrimontal-dreg curves of fihgqxrc10 do not agree Tory Closdy- at-Repciliisnmibers between XJ X 10s and 30 X 20e, the-scmtion drag coef-ficionte obtained oxporimcntally and theoretically aro in goodagroonxmt for Reynolds nuniborsbotwoon 30 X 10G and !jOX 10s. AtReynolds numbers hotmon 20 x 10G and 30 X 106, the higher dra@ of
12 ?LLCAti ““NO IL51
.—
.
—
Beta.mo it WLS poseible to calculati fcr model ~ both the valueof tiaoReynolds number..hb which minimum drag occurred and the vduoof the mction dreg coeflfici,ent.S“athi.@?:Reyntdds numbers, it appw?aw
-.
that i,tis possibl.oto fiyproximwfiethe ~~-s”cdc -effeet curve for asmooth snd fair airfoil by :Ismminflthct.trcnsition occurs at acritical value of 11~ lmtweon WOO md. 90G0 wlon it does not occuras a result of.reversal In tho proaatio gradtont. Because reductionsin the amount oilSurface wavinesa 3rou@% about Iiitle nms.e-arablechen~o in section d.rhgcooi’fi.ciont,the waviness exjutlng on eithermmiel ~ or motLcl8 did zmt P..uzIearto bo suC’iciant@ groat to affect
#
tiiedrag characteristics of therm airfoils ut loc.stat Reynolds
numbcnm between 30 ~ IOG ?md y) x 106. .
NACtL~ No. 11~1
The dxag characteristics of a nmiber ofprr.ctical.-ccxistructlon9irfcll sectiono were
amocth NACA 6Compemd #.Lt
13
seriesthose of
14 NACA ~ No. ~51
f@ure 16 were constructed with chordwise stiffenam. Somewhat~eater dlff’icultymay be experienced in.ccmstructing airfotis withfair contours when spqnwi~e stiffeners Wat are heavy with respectto the airfoil skin are used,
Photographs of model 12 (the NACA 23015 (approx,) airfoilsection) and model 13 (the NACA 23016 .aii&oilsection) arepresented as figures 18 Wd 19, respectlyely. The variation ofsection drag coefficient with Reynolds number for these two modelsis presented in figtie 20 end the waviness measurements are pre-sented in figure 21. —..
The lower bag of the two models wae obtained with model 1.2,which had a section dr~ coefficient of-O.0057 at a Reynolds nurn%erof 20 x 10= (fig. 20). A fair and moothNACA 230-series airfoilwould probably have approximate.y the s- section drag coefficientas model 12, at least up-~o Reynolds numbers of approxina%ly20 x 10=. The waviness existing on model X2 (flg~ 21(a)) in thereglam where laminar flow might ortiarily be expected, that is,up @ approximately 0.12c on the upper s@?ace and 0.20c on tielower surface, evident~v had no adverse Gffects on ‘thedrag of thismodel up to Reynolds numbers of amyoximqtely 20 x 10s. Becausethe waviness characteristics of models 12 end 13 were similar asfar back from the leadin~ edge as approx@mtely 0.40c (figs. 21(~)and .2A(b)),the waves existing cm model.13 in the laminar-flowregion also probably had little effect Orithe drag characteristics*The extreme waviness of model 13 behind the 0.40c position wasprobably due to +Ae very thin skin of this model (tabti). Thoskin was known to.vibrate considerably during the drag teets. Itis possible, therefore, that such vibration was respcrmible for thefact-that model 13 had generally higher wags then model 12.
An example of a model that shows the effect of deviation fromtrue airfoil contour ovema large part of the chord is model k, forwhich drag dati-are presente~ in figure 2P and surface unfairness(deviationfrom true contour) end p~essurs-distributionmeasurementsare pre~ented in i’igure23. The effect .? dmiation froinmntoti-(fig. 23(a)) aa the pressure distribution-was to increase thevelocities Gver the first ~0 percen~hord above the theoreticalvelocities and to move the minti~ yressure point frcm 0.60c toapproximately 0,50c (fig. 23(b)). A comparison of the drag charac-teristics of model 4 with those of model 7 (fig. 22) shows that thedeviations from contour had little effect cm the drag of mcdel 4 at
Reynolds num%ers below 26 x 10s hut at Reynolds nunibersgreater th~26 x 10G tie drag of motil 4 taed b be greater than that ofmodel 7.
.
.
.,,
.
.
.
“
-.
.
NACA ~ No. 1151 15
~~ison of NACA 6- aaii.t?~-sm$es airfoil sect~. - In orderto determine whether the relative merits of airfoil sections ofdiffermt series are masked by construction defects, the dragcharacteristics of several EACA 6- and 230-series airfoil sectionehave been compared..
Drag data are presented tn figure 24 for models 2, 8, 12, and 13.Figurti24(=) shows little difference in the section dr~ coefficic=tsof the NAC.A66(215)-214 (approx.) and 23016 airfoil sectione in theoriginal conditions, although the dreg of the NACA 66(215)-214 (approx.)atifoil section is much lower than that of the NACA 23016 airfoilsection in the finished cond.ithn. Comparison of tho drags of theNACA 66(2x15)-116 and 23015 (approx.) airfoil sections in figure 2k(b)shows appreciable difference in tiag of the models in the originalcondition but a much greater diff6r~ce in the emooti condition. Fromthese data the differaces in drag associated with smooth N.ACA230-and.6-series atrfoil sections, as constructed.,appear to be consiti:=blyreduced if not enttrd-y masked.
Gomw ison of draQ of tiDk Winl?erld.rrractical-construwing model.-
CticnA comparison ha~been~de in figure 25 of the drag
characteristics of a smooth practical-constructionwing model havingthe NACA 66(215)-214 (approx.) airfoil s~ction end a smooth test -lof an airplene wing having tho NACA 66(215)-2(14.7) airfoil section.The airplane wing p.nel had been carefully famed to eliminato anyl?ro~her~c~s or waviness due to wing joints or access-dotis. Boththe airfoils us6d had NACA W-series sectiom With thicknf3ssratios ofapproximate..yO.lh.
Ih figure 25 at section lift coefficients below 0.3, thepractical-constructionw~ng model had lower drag than the airplaneWhg ~iel; whemas,at hi@jr section lift cooffici~~ tie re~sewas trucl. Since data for the airplane wfng were obtained in flight,it is difficult to determine whether tie hi@er drags associated withthe airylane wing wetiedue to buckling under load at the time thatthe data were obt.ainod. It is ~ssiblti, howaver, that w%,~in~ksoi”” ‘-”
—..
the airplane wing existed relati~%ly fm beck on the wing surface,~d the adverse effec.teof such waviness ~re noticeable only at thelcwer section lit% coefficients. Furthermore, similar waviness thatwas not large enough-to cause premature tramgition under the favorablepressure grad.mt exist.i.ngat the law sectioa lift coefficients mighthave existed clcmer to the leading edge of the iTACA66(215)-214 [q?prox.)airfoil section but, undel’a less favora%le presswe gradient at sectionlift coofficientu almm 0.3, euch waviness might well havu roaultedin pr~tur5 transition.
16 IL’,cilm NO. 1151
i“—
.-
—*
—
.
17
mxmtod Wfti-linKim wiIw, which was
23(c) shows that with the model undereli@lt waviness (1.Og) little or no
OffoCt on tlie”drag waElful~.a,but that Wi’& the model utier aload great Gnough to prcduce some perma.nentdeformtfon of tieskin (1.5g) waves e~~stealthat were serious enouf~!to bring &bouta sharp increeue in drag at a Reynolds nmnher d 20 X 10’. ..
For the ceses just considered, SM @t pemwmenli set in the skinor rivet~ of the wi~a caused by compressive loads had little or ns
..—...,_ -.,_. ._._
affect on the drag charactiviat~cs. While the wing was experisnc~load aufficient to [email protected] defmmaticm, however, the dragchar~ctoristlcswera aiiver~ely.affoctadto a considerable e-xtent.
Effects of de.icwg. - Sk+! are presented in fi~ 29 for twoairfoll models equi~?pe~with le~ding-e@ de-Tcer boots. TMso bm)ts
—
calsishi of X“Lb-Df3rS“hee+mat‘%zti.edto the w- sm’f2ca and weretapered to a fine edge cm the uppev .m3 lower suz~aces of the wLr-foil at the point where tLey faired Into the wing cmaxnir.
A O.075c ?e-icer boot on tho leading e~e of IELC&Ill’jj(theNACA 6=5(216)-215 (approx.)Lirroil sectim) cmaod Q soction-drag-coef:icient increment ~mounting to 0.0025 or 0.9030 (fig. ‘@(a)), ‘whereas a similazzO.15c de-icer boot caused increments of appz’o.tinw.tely0.0040. A O.MC d=-icer hoot on Zwdel.U (me NAC~!~~15 (al?mx. )alrf011 secticm) Cause& secti.on-~drag-cooi’flcientincremmts of a.pproxi-matsly 0.0010 (fig. 29(b)). The total section drag coefficients ofthe NACJ.6-ser~es with the O .075c de-icer boot and the N.?IC.423015 air-
foil with the C).LOC airfoil ~e-ico~ boot wero approximately O .~70 ~ReynoMs numberB between l@ X 10s and 32 X 10’, wher~~= tie drmg oftineNACA 6-series airfoil yiith the G.15c de-ice~ boot was somewh&tgreater,Lt leo,6t at Reynolds mmhers up to 3.0X 10=. It wouliiappmr,then, timt not only &r& tile&a@ of .zirfo~1 mctions increasedconsidm-ebly by the adflit.1on of lcadi~-ed@ ~e-tce~ boots but thatthe dificrcnccs in dr~ usually associated ‘wI.W aixfoil sgctfon~ ix!?different-series are ulskod, at least for thtclm-esgratioc of approxi-mato~- 15 pcrccnt.
cowmJsIoNs
~r~l tke analysis of ‘trm dr~ cham.c’terlstZcs of pr.actical-Constzlzctim Winflf3,quaatitativG data Wfirocib~ned that indicatedtho sizo, nunikmr,end locations of surface we,706sufficiont to inducoprmztu.m transition at [email protected] -rnmkmrsgrwater ~ 9 x los, atReynolds numbers greater then 16 X 106, at Reynolds nzudborsgrge~er
Iw.cx.m No. I-lSL
did not Wing ~bout prmkatm.renuwl,ti:’sLm “a Zlmmx.tmateiy
..
..
.
.-
—-
.
—.-—
. ___
b
. .
NACA TN ~0. 1151
characteristics of two wing sections designed to retain true contoursunder loads usually encountered in flight. Whilo the wiw wasexperiencing load sufficient to produce such deformation, however, ““the drag of the wing was considerably higher than the drag of theunloaded wing.
7“ Airfofl sections having thickness ratios of approximately15 percent and equipped with de-icer boots on the leading edge hailsection draG coefficients of6approximate&y 0.b070 over a range ofReynolds number from 10 x 10 to 32 X 10 b Thie value of the sectiondrag coefficient, furthermore, seemed to be indepentint”of the air-foil section upon which the boot V=S mounted.
Langley Memorial Aeronautical LaboratoryNaticmal Adviso~y Comuittee for Aeronautics
Langley Field, Vs., July U., 1946
RE?mRENcEs ————
.
.
,
.
1. Abbott, Ira H., Va Doenhoff, Albert E., and Stivers, Louis S.,Jr.:Sumasry of Airfoil Data. NACAACR No. L5C05, 1945.
—
2. Braslow, Allert L.: fivestlGatfon of Effecw of Various CamouflageP#nn~Aan2 Painting Procedures on the Drag Characteristics of
“(k21)-420, a = 1.0 Airfoil Section. EACA CB No. L&G17,
194h.
3* Tetervin, Neal: A Flthod for the Rapid Estimation of’TurbulentBounda~-?.ayer Thiclumssas for Calculating Profile Drag.NACAACR No. L4G14, 194~.
4. Jacobs, EN., and van Doenhoff. A. E.: Formulae for Use inBoundmy-i.&yer Calculations on Low-Dr&.gWfn@. NACA ACRIAug. 1941.
5“ Davidson, Milton, Houbolt, John C., Rafel, Norman, and Rossman,Carl A.: I?r&limlnaryAerodynamic and Structural Tests Showingthe Effect of Compressive Load on the Fafrness of a Low-DragWing Specimen with Chordwise lh.t-SectionStiffeners.RACA ACR No. 3L02,1943.
, .I
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NACA TN No.” 1151 ““Fig. l_,
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1 iNATIONAL ADVISORY
COMMITTEEFORAERONAUTICSo “
0. .1 .2 ●3 .4 ●5 .6
Chordwise position of transition, x/c
Figure 1.- Calculated variation of section drag coefficient withposition of transition on NACA 66(215)-116 airfoil section.Cz = 0.1; R = 20 X 106.
.-
-.
b
.
,
,
,. I
.
,., , 1
(a) Side bottom view.
Figure 3.- Model of NACA 66(215)-214 (aPPrOX.) Practical-construction air-
~ foil fiection with unpainted surfaces, Model 2.
,.,:, .
., I
I
,.
b .
I
I
“
(b) Front top view.
Figure 3.- Concluded.
, , . “
Figure 4.- Model of NACA 66(215)-116 practical-construction airfoil Bectionwith local surface defects glazed. Model 3*
LhiAL 28028, !, ,!’ ‘, (j
,, !, . ,,
I
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(a) Upper-surface templet.
Figure 5.-[
Model of NACA 66(215)-116 a= 1.0, Cll = 0.2
a = 0.6, Cli = -0.1 1
,practical-comtruction airfoil section. Model ~.
. , , * .
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(b) Lower-surface templet.
Figure 5.- Continued.
I
.
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,
(c) Spanwise variation in contour.
Figure 5.- Concluded..
%’.u-i0.
.
,’
, .
(al Noee templet, model erect.
Figure 6.-[
Model or NACA 66(215)-116 a= 1.O,’ cl = 0.2
a = 0.6, cz~ = -0.11practical-construction airfoil 8ection with Ourfacea ,painted with zinc-chromate primer. Model 5.
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,
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t-
FLgure 6.- Concluded.
,.
1
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. . , * ,
Figure 7.- Model of NACA 66(215)-116 practical-construction airfoil sectionwith surfaces glazed and smooth to rear spar. Model 6. .
0$.-4
1
.
.
$04
a .
Surfaoe oonditlon
Ori.gtial, oanouflage painted
I@tly sanded
Both surfaoea glaaed to 0.120
Upper aurfaoe glazed behind 0.120Lower nurf aoe glaaed to 0.120
Upper eurraoe painted to o.~oLower .mrfaoe painted to O .Uo
Both ourraoeo painted to O.’po
I. .
MAT- ~1-Y~m~m
o 4 8 K 16 20 24 28 32 ?kXReynolds rnxnber, R
(a) Model 1, HA12A65(216)-3(16.5) (aPmx. ) ●ofl -tion~ 02 = 002; te~t~~ m 311 ‘* 324”
F@xre 8.. mreot Or ●urraoe lmprovementa on drag oharacterletios of fiiprOn seotlom.
o
!lDTSurface condition test
Original, unpainted 25
J
Glazed and painted 26Refaired
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comrmEFon MmaAIJTKs
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(b)Hodel 2, I?ACA
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12 1.6 20 * 28 32 36 X ~06~
Reynolds !nnnber, Rc
66wj )-24 hPprox. I
Figure 8.- Continued.
, .
alrfoll aeotion. OZ = 0.1s. El
z
zo.
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*
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surface oondit ion teat
o original, bare+netal skin 192+ fared UP tO EW9r jOtit (o.320) 19El Painted to Bpar joint k
0 @lazed and painted over, apar joint &A Painted all over 202
v Painte# all over; unfair aurtaoea
partly refalred 205
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(o) Model 3, llMA 66h215H16 airroilsection. OL = 0.18.
F@re 8.- Continued.
I
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Surface oontlltlon teat
Orlglml; painted withzino-ohraaate primer 362
Painted 365Glazed 365
NATIW ADVISOQYwMlllEEFmmomJIKs
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Reynolds nunber, R.
106
(d; Model h, NACA 66(215)-116{
a = 1.0,
1
Cq = 0.2airfoil aeotlon. aL = 0.1.
a = 0.6, Ozi = -0.1
F@a’e 8.- Cmtfnued.
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+ Painted end glazed 685
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Reynolde nmber, R
{(e) Model 5, liilOA 66(25)-116 ~ ~ ~:~: ~~~
1
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= -0.1ahfoll. seotion. q = O.L
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(f) Model 6, NACA 66(215)-116 praotfoal-construction airfoil section. cl .0.15 (approx. ).
Figure 8.- Conchxhd.
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. . . .
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.-
,
Figure 9.- Three-quarter front view of~pper Burface of NACA 66( 15)-114alrfoll section in Has-receivedil condition. Model $.
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TDTSurface mndlt ion teat
0. Ae reeelved 898El Faired, both
surraces 912Calculated,
~ = 8000
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o 8 16 24 32 4ol@ 56 x 106
Reynohla nunber, Rs I
Experimental and mlculated seotion drag oharaoteristica forp
Figure 10. -mcA 66(215)-14 praotioal-construotlon airfoil seotiou. cl = 0.1 s
(approxo ). Model ‘7.
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921j❑
Calculated, $: /33:955
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#
6x 10
Reynolds nmnber, R
F@re 13.- Ccmparison of experimental end calculated drag-aoale-effect our’er’elf;rNACA 66(2x15)-116 practioal-construction airfoil eeotion. cl = 0.1. .
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Figure 15, - hlodel of NACA 66,2-115 practical-constructionwith camouflage-painted surfaces. Model 11. (Model 10internal structure. )
.
airfoil sectionhas similar
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Figure 18.- Model of NACA 2301!j (approx. ) practical-construction airfoilsection with camouflage-painted surfaces. Model 12.
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, *.
2“ I. I
Figure 19. - Model of NACA 231)16 practical-construction airfoil section with
camouflage-painted surfaces. Model 13..
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NACA alrfoll o ~~!l! test
aeot ion Model
o a30i5 (aPPrOx. ) 0.30
+ 23016 .15 4$ 12
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}= 0.6, c~i = -0.1 0’1 4 365
Q 66(~5)-~ 0.1 7 9M
r
NATIONAL ~lS43Qy
COMMITTEE m AEawAUllcs
‘o 8 12 16 20 24 28 32
Reynolds nunber, R
(Drag of NACA 66(215)-n6 ~ ~~”~’ Czl}
= 0.2Flgllre 22. -
cl~ = -0.1praOtical-
.saonstructlon alrfoll aecti n, aurfaues- painted and
with drag of falred NACA 68(2~5)-@ practlcti-cona~$!;;;%~!
aeotlon.
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x 106
NACA TN No. 1151 “ Fig. 23a,b
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~.6I I I I I
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.. —-
.
(b) Preaoure distribution, ox ~ O.1.
Fls’ure23.-{
= 1.0,1
Surfaoe uofalrnese end preseure dlatrlbutlon for iTACA66(215 )-116 : = oe6, ~:f ~ -~.~ ‘
.-
praot ioal-eonatruotlon.
airfoileeotlon. Model 4...
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o 23016 Orighll - painted with
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+ 23016 Camouflage painted
: %%1:% 1::$::1 :::~:~:%~:~
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(a) HACA 23016 d
Comphrieon of drag
u? 16 20 a 28 32 36 xl#
ReynoldfJ nmber, R
66(215)-24 (approx,) airfoil sections.
chnructeriatics of come 230- ati 6-series airfoils.
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liAOA ailfoilaeot Ion
Surfaoeaondition
.
q lest model
o 2301 (approx.$ I
Origilml- bare rnetilt
0.302301 (aPPrOX. Camouflage patited and
LTT 192lz
sanded ,33 ~T4. .
66(2.m.j-116: ;
Original- bare metal 0.1066(2xI )-1I.6
L~ 2.05 8(llmedto 0.70 .10 !OIT10
o 4 -0 ).2 ti 20 24 28 32 36 x106
Reynolds ntanber, R
(b) NACA 23015 (appro=. ) and 66(w5)-u6 airfoil seotions. z
Fig ~.- Conoluded.‘~
,
1 g~
.-
Fig. 25 -- NACA TN No; 1151—
30
25cG
oo .1 ●2 ●3 .4 95 .6
Section lift cmeff icient, 03
NAOA airfoil Surface Test Modelraeotlon aon&itlon
ro 66 215)-2(14.7)
,[Smooth Flight
D 66 215)-214 (approx. ) Smooth TDT#3 ;.008 ~ I
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s
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Figure 25. -airplane
.
0 .1 .2 ●3 .4 95 .6Seotion lift aoeffioient, ‘z .
Comparison of drag aharaoteristlasof smQoth teat panel ofwing with that of smooth practical-constructionwing model.
.
.
, . . . . .
(a) Front top view.
Figure 26.- Model of NACA 66(216)-(1s25)16 Practical-construction airfoilsection. Model 14.
!, .1,, .
, , . ,.I
(b) View of model being 8ubjected to compressiveload in l,200,000-pound testing,machine.
Figure 26.-Concluded,
I
*
(a) Front top view.
Figure 27.- Model of NACA 65(216)-215 (approx. ) practical-construction airfoil section. M~del 15.
‘$:Ip
* 9● ☛
. .
Figure 27. - Concluded.
,
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o Before loading?$
w
After loadingP
+U+
NATIONAL ~
ulNNITm Fa#mnMI’Tics. —
o 4“8 12 16 20 24 28 32
Reynolde nunber, R
(a) IWA” 66(215)-(1.25)16 airfoil seotlon with ohordtise hat-seotion stiffener npaced
3 Inches on oentera. o~ = O.~. Hodel 9. .
4 E’l@rO 20.- Etfect of oaupreaaive load on drag oharaoteristica of alrfolln.
,,,
36KL06
I
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..
%urfaoe m . .Ool’m tion test
o Ihfore loading 389+ ~ter l-tig 397
o 4 8 12 16 20 24 28 32 36 x 106 h
Reynolds nuaber, R *
(b) IIACA 66(215)-(1.25)16 d.rfoil seotion titi ~H~d~y~hat-section stMfener@Elz
spaced 6 i.nchea on oenters. c1 = 0.16. .z
. . . . .I
.
●
Surfaoe oondit ion
c1 Unloaded+ Loaded sufficiently to pro-
duce sllght waviness•1 Loaded mlflclently to pro-
duce pemmnent deformation.012
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(o) NACA 65(216)-215 (approx.) airfoil section. OJ= 0.3. Teat, m 38.
Model 15.
Mgure 28.- Conoluded.
I,,
“1
106
o 0.30 smooth m~ 2+ .30 0.0750de-~oer mm. 16
}
15❑ .30 0.150de-leer LT1’2~
Reynolds nuuber, R
(a) 0.075c and 0.150 de-icer boots on i’iACA 65(216)-215 (approx. ) airfoil seetlon.
FigW 29.. Effect of de-leer boots on drag oharaoteriatics of airfoil sections.
.
106
m . . . . .
1 . ,
.1 ,
● (H2
.008
● 004
0
c t Surf aoe condition TDTtest Model
o O*33 Smooth 4●33 O.1OC de-icer 4 12
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—+ —~ .+ +
/ wm
NATIONALADVISORYCOMTTEE m ~lcs.
I Io 4 8 1.2 16 20 * 28 32 36 x
Reynolds mmber, R
(b) 0.106 de-icer boot on NACA23015 (approx. ) airfoil section.
Figure 29.. Concltiecl.
106
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