Document
I
RM No. L7I22 ENClj",,EFr'.%G DEPT L'' }y CH A IRCRAFT .
N ' . N.
A
RESEARCH MEMORANDUM
AERODYNAMIC Cll "CTE ISTI OF IFIL+'D NACA FOUR-DI U
anc; k4V* S. Cohen
h )aiical Laboratory la., Va.
NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
r:IASHiNGTON HACA RM No. L (I22 NATIONAI. ADVISORY CcH•tI'I'_IM , FOR A.EE,,O11 IWICS RE M "ARCH AUWRPCMWIi A ROM IC CRA'2AGMISTICS OF A Pte' Or , 140DIFIED NACA FOUR -DIGIT -. 2 -KEMIEwS AIRFOIL M 1 IONS By ia.urence K. Loftin, Jr. end. Eonneth S. Cohen `I'hooreticrl -oresstara distributions have bean :calculated and the
eaparimntal aorodyn ,dc charactorlatics d.a'Wrr>_ wcl at low epeads for
a solected soup of the NAC.4 Pour-digit-aeriao airfoil sections. which haA previously baon modified for hiE,h-speod application3 . The experinontal investio.tlon which was male in tho Lan g ley two-diMen3ional low-tu?^a v a. on ' ce presBure . 1 = , .ol conjisted of ?^ , xeazuremc nts of the lift, drag, a_A itching-nomant characteristica of each of the plain airfoils p at Reynolds umbers of 3.0 x 10 0 , 6.0 x 10 6 , and 9. 0 x 10 6 . In addition., the ©ffectivenaes of flaps whoa a p plied, to these airfoils and the e+fect ix-,on the aerodynamic characteristics of standard^leaddne-edry roughness were determined at a Reynolds number of 6.0 x 10 0 . Also tasted ware th r ee conventional TEM four -diCit-sories airfoil sections which had not previously been invoatigatad in the Lang two- dirne naloiial low-turbulence pressure tunnel.
Tho results of the ex-perizental investiption indicated that the lift characteristics of the modified EI i 1CA four-di^i.t-series sections having normal.-sizo loa iri g -adgo m ii ana a maxfmom thickness ,-ercant of 12 chord locatod at 40 percent chord vex-y closol;r approzinated those of siwoth 'RAC-A 64-ser-'es low-drab, sections of correspondin- thicimass and.- camber. than the 1eadin[ -odd radius was reduv3d to one- a uar ar nor al size, the maxim= lift coefficients of the 10-percent-thick mxi_laum airfoil; with r- thic!cwss lovated at 40 and 50 percent chord ware about 35 percent lower than those of 11AC-A 64- 3ories sections of corre- aponding , tnic mess and camber. For airfoils e3quip;ed with 20-percent-chord
m axlm =
split fla p s deflected 60 0 , the lift of the airfoils with one- cuartor normal-size 13ad_ TW-edE1a ra3.ii more noarl,-T approached that of NACA 6 , 4-serioa airfoils. Rousbwss had no a p preciable effect upon the t axi n= lift of those airf oils . The min immz drag coef ficIonto of the -aximu airfoils with t_hicime3s at 40 percent chord and normal-size loafing-edqp radii ward higher than those of the correopondin` NACA 64- :series sections. Reducing the leaii.ng-edC p radius to one-quarter normal size and moving the position of maximum t Malmo sa to 40 and 50 percent chord caused the minirnarw_ drat eoofficients to be reduced to
NACA ^ >Io. L-722
056- aeries
values about the as tho;a of corresponding uICA 614- and
a*,1e resulting
-actions, reopectivel;y.. Increases in the trailing-ed_z
from roarwsrd m:ovemnt of the position of m tim m thickness caused
sharp decreases in the lift-curve slope and ixronounced forward move-
TiP_iic center.
mant of tha aerod;, ,
l NTPOMICTION
The increasing demand for Mob speeds is md-orn airplanes has
p eration at
focused much attantior_ upon airfoil sections Canabla of o
the adverse effects of eom p roseibilik .
high IAach numbers without suffering ,
One of the first systematic series of airfoil sections develolmd With
a viaw toward high-spee& application. consisted of modlftud XACA four-
digit-series Bastions. Descriptions and him Mach nmber data obtairaed
in the 1-inch hida-spoed tmnel were presented in 1934 (reference 1)
for those air-foil sections. Since the issuance of reference 1, the
modl.fied IMCA :four-digit-series sections have bean u=lo;aed rather
extonoively in Europe, p articularly in Ger i and have recently
received favorable consideration in this country.
Low-s )oed aeroct:; r namie data obtained in the 1RCA Variable -Danaity
I
Wind. Tunnel are available for sevoral of the modifiad IL CA four-digit-
of
series airfoil sections (reference 2) . Tha r=7,e airfoil types
view of the meager
coveroi by these data, however, is very li=ltsd. In
mount of da t a available for the modifies HACA four-digit- series
;sections and because of the recent int -est shown in them, an invasti -
ention of -Uie low-speed aerod;rx aaic characteristics of a selected group
was x dartaL-an in the Laney two-dimaneional low-turbulence pressure
tunnel. Tha airfoils choson for test were those which appeared from
theorotical pressure-distribution calculations to offer the best
possibilities for hi Ch-speed applications. The results of the ex p eri-
,ntal i aaeti tion, together Frith the thaorotical pressure-distribution
data. for a. i I CA four-digit-series sections, are
y .=bar of the modified NA
presented in this pa,-xer.
The aerokaia l sic characteristics of five of the m:od'. ied sectiorv3
are presented; three of thaws are symnstrical and two are cambered
with the NACA mean lire a - 0.8 (modified). (Soe reference 3.) Also
presented are characteristics of tr=ue conventi3nal NACA four- d-igit-G-3rie8-Bec'1iOnSY
data for which are not included in tha esystomatic
results of reference 4 for thin aorias.
CO g 'FZCI - U .7 1 AIm SM,IBOLS
ed ;section d *TIL T coefficient
cdinin ' minim= section drag coefficient
IUCA ?U^I No . L7I
22 3
C Bcction lift coaf icient T`SaxiL
C3 = section lift coefficient
c ti desiga section lift ooefficient
section pitching-momnt coefficient about aero4 r na is center
^ti'lsG.
section pitching-monsnt cosfficient about gmrter-chord point 0%/4
ao section angle of attack
section ale of attack corresponding to design lift
coefficient
dc I Pjx
section lift-curve slopa
V free-etxeam relocity
V local Volocity
AV increra nt l ocal
of' velocity
a
Ov incremQnt of local velocity corresporAing to additional type of load distribution
resultsut p ressu r
e coef^icient; differ©nco betvman local RR
upper - surface and lover-surface presam-o coefficients
R Reynolds number
R5 boundary-layer Reynolds number based on botm ary -layer
thickness and local velocity outside the boundza7 lajar
c airfoil chord length
a
diotance along chorl from leadW., edEp
distance perpendicular to chord
V
yc mean-lira ordinate
a
man-lino designation, fraction of chord frc leading edge
over which design load is uni_fortn
DESCRIPTION AIM MUMETICAL C'MRACTMISTICS OF AIRFOIL.,
Banec thicImess -Lox=.- The nod.i ication to tho NACA
foixr-dit.t-
series basic thicknase farms, completely described in reference 1, can
par}tapa be bast described hare by an explarvation of the diCIt, al)pea-riag
4 PIACA R1 No . L7122
in a typical airfoil deli ,nation. Consid er, for axample, the RACA 0012 -64
attached airfoil section. The fir3 t fotx digits havo the usual m^-f3ninE, to the numbers arfaearl.nL! in the designation of a conventional :ACA fou— _ digit-series airfoil sacticin, in this case a 1^--?ercant-trick eymmtrical section. TI° , two numbers following the da-ah doecribe the modifications.
The first rumbor followi.no. , the dash is an indox to the size of
loading-3dc radius. LaadisiL-Qd radii of t'ravo sizes, repres®sled
the
by the numbers 3, 6 find 9, were investit^ated in reference 1. The
number 6 which arpears in the illustrative oxaspla indicates the norral- sizo loading-edge radius employed with c;,nvantioral four-digit-series cocoon- -;the nrmwer reprU3en_ ,s a one-quarter r_ox%:sal-zizo 1eadir4_?-edLe radius; wul the number 9 indicates a leading-edg ge radius of three time,
normal size. rho second niu•iber following the dash indicates the
maxImlim thic cress in t-3nths of the chord. Airfoils,
position of
maxianua thiclaiess
which were darivocl in reference 1, have the positioni of
40, 50, and 00 percei;t chord.
:iccated. at
In order to urovide some basis u p on which to choose the airfoils
p
to be ' asted, theor3tical pressure distributions were calculated by
the methods Gf refer©nco 5 for a croup of modified NACA four-d git-
series basic thickness forms. The recruits of these calculations are
presented In fio rea 1 to 8 for -the followinE, ail-"oil sections:
ICI:CA 0010 -64
RAC-A. 0012 -64 LIACA 0010- 65 IVY GA 0010-66
race 0008-34
IMCA 0010-34 I•Ir1CA 0012-34
HAGA 0010-35
In addition to p ressure distributions at zarc li-^t, throe data include
' incremental velocity ratios from which the prossure 0—I stribution at ar;I
lift coefficient nay be calculated. The method of making this calculation
is described in n3forence 4.
From - the data of figures 1 to 8, the effect
u p on the pre s sure
distribution of vuriaticns in the position of maxi=m Chic , 4eaa and
size of the lead's-edea radius are clearly evident. A decrease in
both the peak ne ; tine pressure coefficient and in the variations of
pressure over the forward part of the airfoil is e_"f octed by maintaining
a normal-size laadino-edge radius and novinZ,
the position of T'x 1no n
thickness from 30 (original position) to 40 rorcent chord (fig. 2).
Ilarther re rd movement of the position of naxirun thickwes, ho-imver,
appears to cause a secon& peak in the pressure distribution near the
tsailinS odrre (firs. 3 and 4) follovod b,-; W
railher shar2, undesirable
p reasurc recovery. With ow-quarter norml-size loaiinC-ed,ro radius,
the ma g
nitude of the peak m,-_tivo pressure coe:^f_icient is not cYAnged
such but its p osition is moved to the rear. The chan=ge is position of
nrdsE; - ,xra is narticularl;; marked whon the ^sosit
m1nim-m ion tf maxis,,=
thickliess is moved f rora 40 percent to 50 percent of the chord (figs. 5
XACA P14 No. L'a22 ,., =
=ax:! thief meos decreases
and 6) . 'Phis mowLi3nt of the position of
the p eak na a.tive pressure coefficient sli;--htly but results in an
undeairabl z v lard pressura recovery near the trailiz^ edge . On the
basis of these theoretical data and from a ecnel.daration of the prcbable
low-speed chars.6 toristics, the IUCA 0010-34 4 . 0010-35, and 0012-64 basic
i 0010-34 aad Ool2 -64
thicImeae for...s umra chosen for teats. 'fhe NAC1 1
„rare also tested iz combination with a caribored Lie= l na- -
Mean line . - In the ;oret3ent investigation, tha modified. MCA four-
dif^lt-seriao basic thickness forte which were cambered employed the
IACA a = 0.8 (modified) mean lire (reference 3). This meam line is
dasig -iod to have a uiiifora load diotribution f rom tho leading ed0a to
the " -b-percent-chord station and deai L -.;ned to be Goo:flatrically straiCht
froaa about 85 percon't chord to the trailing edge. The ::ACA a = 0.8
(mod.i led) ::jean line wao u ed because the yak ine-laced velocities added.
by this mean line, to the velocities over the basic thic3awss form are
less than those aGsoeiated with the older wan lines, such as the
ARQ I G 230 ani 24 man line; and the curvature of the airfoil surfaces
near the trailix , edCp which results from the use of an MAC."i a - 1.0
mean line is elimir ted.
Ordinatez and load-distribution data corresponding to a. design
lift coefficient of 1.0 are preoented in fiLmro for the MCA a = 0.8
(modified) mein line. if the ord=inates aua loan arcs desired for a
design lift coefficient other than 1.0. they may be obtained easily
by liuearl y scalinZ the values presented. The method for combining the
preosu e-distribution data for the basic thi.cimess forme and mesh line
to give the pressure d-istribution about a caaberod airfoil at SxVj lift
coefficient is given in reference 4.
DesiU,nation of , ca.mbered airfoil sections.- iuo method of clesiLgi ting
modified W-GA four-dieit-series airfoil soctionc which omploy the
MkCA a - 0.8 (modified) mean lire is illuotrated by the folloving; exam:,le : NACA 00122-6^ a = 0.81 (mcdi:iea}, c ti z 0.2
This system of n7zbers clesigz t.es an MCA 0012-64 basic thick - ness fora
laid off on an IMCr a = 0,3 (modified) neeL r
i line cambered for a
daaiga lift coefficient of 0.2.
Conventional 'NACA four-digit- aerie ; airfoil ions.- Complete
sect
descriptions of the basic thic^ss forms and mean lines of the
cc:ventional TIACA.:our -digit -serial airfoil sectionU of which three
imra tasted in the present investigation_ may be round in references 4
and 6.
NACA M4 No. L71212 1F4&TS APPILT.'.ATM AND inve stigztion wao made in the Wind. tunnel . - The Langley t-do-dim noicnal low-turbulence press •;zre tunnel. The test section of this tunnel measures 3 feat b. r"•5 feet wi Lh the models, when momted, ct.^mi;letaly spanning the -foot IL.wnsion and with the Jmcture ee . Lift between the model end. tunnel walls sealed to prevent air leakae neasti re;,,ents were made b-,,, takdn , the difference between the pressure
measurements
reaction .apon the floor and ceiling; of the tumre1, drag.
were wade by the wake-stove;;, method, and pitchinC moments were daterminod with a torque balance. A more complete description of -tie twtu_ ,el and the metl4ods o: obtaining and reducing; the data are contained in reference 7.
airfoil sections for which the a peri=ntal
Modelo.- The eiijA
e.erodynaxmic chvractorietica were obtained axe: oC 10 -3; NACA NAGA 0010-?4 IIAC A 0010-34, a - 0.8 (nodifiod), c ` 1 = 0.2 MCA 0012
w,GA. oo 12-64
& 0.8(modified), c Z 0. 2 iTAGA 0012 -64, ziAGA 2 40 IIACA 2410 The modela repreoenting the airfoil sections Vero of 24-inch chord a d, with the exception of the u--percent-thick section i&lch was machined froze steel, were constructed. of 1&,TInated mahoearV. The models were spra^-el with lacquer and then sanded with Iio. 400 carboru=! *n paper until aerod;, nandcally amooth suz^ aceo Fr3re obtained. The ordinates of the modela tasted are presented in table I.
The tests of each smooth airfoil eactionn consisted of
Te sts.-
measurermnto of tha lift, dx 4C,, and gran yor-chord. pitching nomant at Rew-noldc numbers of 3.6 x 10o , 6.0 x 100 , and 9.0 x 100 . In a,,Uition,
wex x detarmined at a
the lift ard. dre` characterigtics of each eaction e
Reynolds ni=ber of .0 x loo with 3tandaxd rouj:IU.I6ss applied co t;r_e leadinc ed^,ia of the niodal. The standard roudiness a=lo--ed on these 24-inch-chord *_ odela consisted. of 0.011-inch-di.a Teter carborundim wins enroed over a surface lenLnth of percent of the chord back: from the laad ^- ed,:?3 Q*1 -the upper and lower surf'acoc. The L-xains were thinly
spread to cover from 5 to 10 percent of this area. In an effort to
apDlied to these gala sotw ides, of the ef'feetiveneas of flaps wher.
NACA FM 17o. L (322 7
airfoils, each airfoil was fitted with a 0.20c simulated split flan
d.eflecte" 00 . Lift measuromonts Caere made at a - Reynolds n rrfaer of
6.o i both Y. loo with the split flap , ie th the airfoil loading. ed, smooth and rough The resulta obtained from teats of the eight airf oil sections are presented (?igs. 10 to 17) as plots of ateadard aerodjnaryic - coefficients representir43 the lift, dra`, and quarter-chord pitchinp -monontcharacteristicsoftheairfoilsections.Thepositionofthe ^ d from the erpernonta.l results, aarodynanic cantor, as daterr-j and the variation of the pitching'-momnt coefficient, about thin point are aluo included. The inflwnce of the tunnel boundarioa has been removed from. all -the aerodyTien is data by manz, of tha fol I cwinG equations (develo p ed in refercanc3 7): c d = 0.990 cd' cz = 0.973 cZ' /4 = 0.951 cz,-. '3%/41 = 1.015 0&0 a o where the primed quantities ropresont the -moa.surod coefficients.
DI:SCLESION The discussion is primarily concernad. with an anlysis of the effects, as -shown by teats of the five zuodi2ied UACA , four-digit-series airfoil section.:, of variation.: in the leauinC,-edge radius and position of maxi== thicimess up on the aerod,aAmic che 'x-acteristics. In this am lysis, frequent use is made of cross plots (figs. 12 to 21) showing the characteristics of the f miodi iacl sectlor.:3 as compared with those of
the corvontiona,l pROA four-di it,-j er:ies section, an3 EACA 6-series
low-draS sectia s. The comspara.tive results fcr the X1CA 6-
series and four-digit-series sections are shown in the form of curves, representing faired data ta1mn from reference 4, whoroas -the results of the present K t V f3 NACA PM No . L7I22 p erimental points.
invPatigation aapear in the cross plots as ex
tha three conventional
for
mention is rada of the results obtained
Little
as they follow closely
NACA four-d.iLdt-series sections tested inasmuch
4 for this series of airfoil sections.
the trends indicatod IL reforon ce
DraC,
Minimum drab.- The previously montionod influence upon the pressure
gradients over the forward part of the airfoil of a :cod°.xet ion in size
ed,7a radiuz aana a roarward movemont of tha position of
of tha leadin` -
favorable effect upon
ZaX iMum thictaleas Las, as might be expected, a
An indication of tho
the value of the minimum drag coefficient.
magnitude of this effect may be paimd froze f iLr,.ara 18, which shows
oection,drag coefficient corruspond.in G- to a. Reynolds
the minimm
number of 6.0 x 106 as a function of airfoil t}aicic oss ratio for the
five modif iod HACA four-digit-series airf oils, for the conventional
DIAC,A four-digit series, and for -the MCA o4- and 66-series low-drag
airfoilo .
,r.ir ^_ I-
mum the 10-porcent-thick
In tha smooth condition, the drab of
airfoils ha,vin, laadinE-edoa rani of one-quarter normal size and
axd 50 percent chord eras of the same order,
ma:cimum thic^ss at 40
66- aariea low-drat
reapectively, a;, that obtainad for NACA 64- and
airf oil,3 of comparable thicknoze. This similarity in drag ind:icatea
lami nas flow over the airfoil surfaces.
the existence of considerable
The s 11, though rather oxtebnsive, positive r^ressuro gcadiant, which
of the 12 -percent-thick airfoils having
occurs over the e au aces
maxim= Imes
-ed.ga radii of normal size and thic s at 40 r/arcent
1ea d.i.ntg
chord, gives riso to a mininrcrm drag coeffici©nt which liras betwean
- :aerios 1G-w-arag section € nd. HACA four-diezit-series
thoso of tho ITACA 64
section of cormarable thielmess. The add_ti6n o° the RACA a = 0.
, - basic thickness
(modified} mean lima to tha INAACA 0010-34 ana 0012 64
drag
fo rms does not appreciably effect the value of the mini
coefficient. The faired data of ref oreuce 4. Valch a-a nresonted
L", figure 1:3 1 indicate that airfoil thiclneas f orm and man lim haTe3
min =am
little effect upon the value of the drag coefficient when the
airfoil leading al oo s are in the s: oaCII condition; and the re ou.l.ts of
the present investigation (fiZ. 18) follow the sam trend.
The airfoil basic thicIneso distributionz a pp ears to have a
Lexkad e2f l ect trpon the man-nor in which do iaini._-=, drag coefficient
varies with lieynolds number , (figs. 10 to 14) . The controlling action
of the airfoil pre3o . ure distribution upon the extent to which the opposite
effects of a thinning boundary 1a;; er and a f orwars movemont of the
-point of transition balance each other as the :naynoldis number is increased
su;gesta itsal' aU a possible ex?lanation. Boma iwight into the
which the airfoil reo3ure distribution influences the
mochemism by p
movement of tha transition point with Reynolds number iaay be gainod
U - 1
from the theoratical work. of Schl.ichtine; and -ich (reference 8) ,
thi. work show tia existence of a critical boundary-
The results of
layer Raynolds n=-bor abova which the la• iixar boundary,
R5crit
NACA iRM Wo . L710 -2 9 layer is no long stable and may become tvrbolant. rurtki3rmore, the va.:Lue of the critical bourZ ..a 7 -18; er Fo7nold3 n1wrl)er is shown to dacreaso ra pill4 - and.'tha laminar boundary l..ayor to become increasi=z17 unstable as the pressure ---radian along the surface becomes positive.
In the presence of an unfavorable preesunra gradient, the transition point is theref ore, most li2aly to move rapier, forward once tho critical boundary-layer Re,^uoldz n'mbar has been reached.
In consideration of tho idaas of Schlicting and Ulrich in rolat.ion to the increase of minimum drab, with noyn.oldo number shown by the the unfavorable ?nresaux-o eradient w^CA ooi2-64 section (fit . 13), bo roeponniblo "or a rapicl over this airfoil (fig. .8) wou1R seem to forward movemmnt of transl tion which overbalances the nornal thiruYinp, of the bowxlay r layer and conaeovant .redue ;;son in dreg that us,„a.ily y o accompagt an increaea in Reynolds number. On th© other hand, th I.= d010 -34 (fi . 10) and I'•',ACA 0010-35 (fig. 17) airfoils w2h ch possess mzcre favorable p-raasure gradients have a naglig ble scale e ct bo twoon Reynolds nursers of 3.0 x 10`-) and 9.0 x lou . ih' s ffe fact' indicates that the opposite effects of a thinnJl-z^-, botud.ary layer and a forward mov 3 iriant of transition nearly countoroalance each other.
-. of WA. 5-serieq The uuiforriLy favorable i flucricu upon the rain:=An d.rrai sections of increa3ing the Reynolds nurlbar from 3.0 'x 10o to 9.0 x 100 indicator t1Mt PE) clit of t--aae airfoil sections, which have misrI:ad ne,ga 'ive gradilawts, is su^^ficS.en'aly hi gh so that no appreciable forward muveL-Bnt of 'Lransi tion occurs between the3o Rey nolda numbers; arid, thus, the favorable effect of e. thinning boland xy layer -predominates.
Low-da-ag ran,- ,,j . - The ran,-.e of lift coeff icients over which lour: drag- is obtained and the :^^ia =r in which this ranE varies with Rejuolds
number are about- tlio saw for tha ?'IACA 0010 -34 and 0010-35 airfoil
sectiona (figs. 10 an;: 12) as for the IVA.CA 6-jorias sections of comparable thic mq ss (rofurenca 4) . The low drag; ra;n p for tho imC11. 0012 -64 Uecti (fig. 13), ho;Araver, is q_u.ite X11 at a 2aynolds
number of 3.0 x 100 , and is practically nonexistent at a Re,!nolds
number of 9.0 x 10u . The more, positive pro sauru -radiont3 on Qie I'ACA 0012 -64 section aro probably responsible for the behavior of tle low drag ranee on this airfoil s©ction.
Tha relationshi p be tzfoan tha drag and lift outside they lour drag range of lift coefficients is about the aaae for the "ri? CA 0010-14 and NACA 001?
-64 airfoils, both cambered and Luncamb3rod, as for the
I+iACA 64-seried low-drag sections af comparable thicinossi a somwhat lass marmad correspondence ex.iats betwoe.0 the drat; characteristics of the 11AC.A 0010-35 section and a comparable NA I CA 6rl-serios low-drag section. These com:oarisons aro valid for the airfoils in both the smooth and rough conditions.
10 Ii4CA RM No. L7122
Lif t
Lift- curve slo-oQ . - 2earcrard. movemnt of the msition of maxim,, .
th3.clmaes of tho NUCA four-dioi .-seriva sectic:is is accomaniad b:; an
increa$3 i n angle. In a.cc:)rdance with nroviouo exia;rinental.
work (re- f orences 9 and 10), the lift-curv3 clo--:e dooroaaes with increasing
trailing - ed,Ee an`;le . The restclt.s of th,: prasent investi L , , Ation (fi t s. 19)
for the 10- ?x-orcont -thicLarA 12-percent-Uhic y sections haviaE max
e
thickmas at various positions indicate th saam trend, rvrith the greatest
decroaze in t-he lift-cvxwe 310- p e boirE. about 1^ percent.
Prom theoretics) consideratims, the lift-carve sio_ne should
incros.ae with increasing airfoil thic— 'aless ratioi, anc. the co^narativu
p rom roforenc;; 4 (fig. 19) for W.CA 64- aorias low-drag sootions,
data
which save v:r; , mall trailing-ad.Ee angles, inclicata that such is the
Cace. Tf, howover, the crailin^?-od^n is lases aaZci incnaasos
e
the theoretical
rapitllj with increasing airfoil thic1nos3 ratio, increase in lif'-,-curve $lop3 with thicim as will be ovarbalaneod by
an^,lo . The NACA
the o p posite effoc t of incmacin.e trailing-edg y
four-digit series sections, da^a for which cre prasontad in fiL.ure 19,
have this cheracterictic3. Since, with inncroaoin: thicl -Was, tho trailin:ry-
oaro ar ;los of the modified i' UA four-di3i.c•-sarias sections become
progres.317aly larCpr th.. r l thcsa of thi convonti-mal MOIA four-d.i,-,it-sories
se- ctions, a mora ra-, p id decroase in lift-c=c. 31ore with increasing; thic.', esa would be eraected for those m:)dified airfoils. The rmu)unt;
of data available for the rod.ifiod I NACA four-digit-series sections does
not ap_peza
r to be suf f iciont, howover, to define adoquat©ly- this trend
or to pormit amr de init,^ state mnts as to the raia.ti7a effects of
,hneas on `rho lift-curve slo p es of the modiTiod and conventionFa.l
3ACA Your-di 1t-e;orias sections.
IXAC
Angle of ero list.- here axcears to ba no mx)rociable difference
In the section angles of zero lilt of the MCA 0010- 4 and NACA 0012-64
airfoil soctiona car1er©d with the W.CA a - p.0 (modified) groan
line (fiO3. 11 and 14) . The values a r -a sli;",htlty :core na_gative than
those -oredictod From the tbaorutical man-line data pre&entad in
fiur g bat agree quid: Wall with values obtaine&
for cambored. 1 7MCA 6,1-serios airfoil sections emlo;,-ing the UA.CA a - 0.8
(modif ied) nx^an, line (reforence ;) .
upon the rja„Yimam lift
lolaxirmmi lift.- Sow idea of the effect
. a coefficient of variations of the Do3ition of maximum , thiclmess and
leadisig-odge radius maj be mimed from f Cure 20. T4is fi,uro shows
Wia maximm soctior lift coefficients (H = 6).0 x 10 0 ) for ,he
LACA roar-digit-series airfoils as a. function of airfoil
m:difiod
.tic, with co:,ma^tivo d- to from roferenco 4 for
thicim-: as ra
NAU, 64-series 13w-drag airf oi13 . As rsiCht be ox ipactad frrom
- = lift coefficients
rrovioua investigation, +lie loirost rnxii:l
w:-)r- obtained for the airfoils raving one-quarter no_x-mal-size
viaxi-- m lift coef ficiouts of the
loading,-edge radii . The
e " '' '^ NACA RN1 o . L7122 ' 11
two s^T:metrieal sections ( 19A 0010 -34 an- 0010-3F) 8M about the saw
and do not appear to vary as the leading,-odr y condition is chan,3od
smooth to route • :.'hose ros-aits show that if the leading edge
from , usual i v)ortc.*_it iai luence of surface condition
Is auf'f iciontl, ; sha y _,, the
L7ELXi,` m lift obtained is negligible. The extreziely low value of the under these conditions is shown by comparison with results for the f t coeff'iciento of tho two modif iel NACA 64-010 section. The oz-Liram li HACA four-digit-:;eries E3ectiona axe about 35 nercant lower than that 64 -010 section in the smooth condition and about 15 p`:rcont of the IM CA lower dh3n the leading edges of tho airfoils are rouji. `Psa incroment - - 34 section io about in mz izim lift cauzad by ceadbering the KACA 0010
sa-a as that obsarved for the adcUtion of a prozimtely the same
the Even with camber, the mmo n-t of camber to the NACA 6 4 -010 oaction.
imma lift of the ACA 0010-34 section is about 23 percent lower raax than that of the 'mkcA 64 -010 as e ti on; but iri th roug,,4 leading a dca , the IUCA 64-010 section has a ma-zi lift coefficient .,rhich is about the 3a a as that of the cambered NACA 0010-34 section.
The maxim= lift of tho three airfoils having, one-Quarter normal- size loadin, -ed.ge radii with snooth leading edrpo ard. _ equipped with 0.20c split f'lap3 3©fleeted ^^ p°^ more marl;= approaches that of MCA 64-series low-Ime. , sections of correopond.in` thickness and camber.
The do ,3xv mont i : maxim= lift co3ff icient caused by leading-odge.
rou.ghneus is, however, so small for these three sections that in maximum lift Of tka .throu wd.'ified.
Lhe ro-a h ccrJ_iUon th:; MACA four-ddgit-series sections is as good as or better than that of corroo-omd.ing ITACA 64-caries airfoiU - Moving the position of mximum UhicYaiass from '30 percent to 4J Wrcezxt rr}Al chord while maintaining a no -size leadinr -adja • r ad_u reduces the d rum lift coefficient of the plain airfoil about 15 percent, as shown by the comparative dales for the TACA 0012 & T'ZACA 0012 64 - sactionc.
Claarly illustralyad here is the important point that a reduction in thickness of the airfoil near the laadiag ed.Ce s , such as occurred in thin caae, has a definitely adverse effect upon. -the naxi zLn lift coefficient although the loading-edr,.a radius iteolf n y not bo docreased- y The L'3 x l ma l lift coofricients of the cambered P!r-1 uymmet rica,l IMA 0012-64 airfoil sections in froth the anwoth condition and with etand rd leading-odga roaC^hwss are noarly tho one as those of the correspondinG cambered and By,musical MCA 64-serios low-drag sections (fig. 20) .
IrAri value of. the mx
irtnn lift coef"f icient presented in figm-o 20 for the imcjl 1 001=-64 section is about 13 percent lowor than that indicated by teets of th© a air-foil i.-,. the YMCA Vari able -Dena ity Wind TL S l (roferenoae 12). `fho value o1ota.ined in the present invasti- ention, hovuver, was vcr,-1 carafull; , checlml al-d is believed to be correct. I-Miej d.iacrepancy otwaen the va!Lw6 obtainod in the two tunnels may possibly have bean caused by turbulonce affects not fully accounted for on this 3enaitivo airfoil by the effective Reynolds t=ber correction anpliad to the Variable-Densit,- Wind-^_ulnal -rocults.
IL&CA R4 No . L'7I22 The results presented in fi`rura 20 show that-, in the amooth condition at least, the rr-xitmam lift coefficients of tL-© cambered atU q
trical I IACt1 001% -64 airfoil stectdcas, than e uipped with 0.20c
symae 1 0 are sowulat hirer than thoae3 of corre- Split flats deflected 60 ,
Qponding WCA 4- series sections. This result Lzr be explained by
airfoil is
the fact that the trailinj-edge angle of the MCA 0012 - 64
larger than that of tho IUCA ' -012 aix-foil since the experi:x:ntal
results nrosent©d in reference 3 Indicate a alight improvamont in _mxiTwo
o-c,erias sections with Sp lit flaps when tine
the lift of I NACA '
ant
trai ling-ole p cusp is reinoved.. Tho resents for the cambered
symmetrical IIACA Ofl12-64 airfoil with. rough leadl iq,- edrros do not farm a consistent eoaarison with results for the IMCA 64-caries oec ions , T-11 noithor caGe however, is the modified. NA.CA :our-digit-3erices section NA GA 64-serie3 a=rcoil.
worse than the correspoudi->6, r Batwean 3 Holds mmbers of 3.0 x 100 and 9.0 x 100 , nom of the raodified XACA four-digit-series sections show aiV appreciable scale effect on 7aazi nm m lilt.
P it ching Momnt; Quarter-chord -ooint.- he two airfoils cambered with the TIACA e - 0.8 (modified) man lire have quarter-chord pitching closoly smith those predicted mownts (figs. 11 and 14) which aGreo from tha theoretical pitching-moment data (fig. 9) .
Aerod,-m&zic canter. -
The ehordwiso position of the aerodr.a-nic
center for tha :codified IuACA four-digit-series sections is shown in
figure 21 as a function of,airfoil t ^3clmacs ratio, tometiier with
►
aimilar data taken from reference 4 for the conventional IIACA
four-digit-series sections an-d the ;;ACA 64-sarieo low-drag sections.
Mie forward =vem i t of the aerodynamic eentar iv!Uch is :eean to
aeeoma4r rearward mvenont of the position of zexinvm thiclalesc on
the ;,codified NACA four-digit-series sections is in afxeewnt with
the trends of ref 3 renco
11 which show that such a forward mova.ment follows an increace in trailing-©dej axiE;7o . T.F:oorotical considerations indicate a roaanrard move-wont of the acrok-i-Anncic cantor With increasing airfo" tliiciwas ratio, and the data for IMCA 54- sorias sections fell--.r Chia trendi but the offoct of inereasinC.. -^rai.iing-ei ana -3
predominates in- trio Case of the conventional. EI rCA four - - aerie s
cliCit
sec tiona a3 evidenced b7 the fo-m and movownt of the aerodjnmidc center.
fig. 20.) $i-vae the trailin, See , 7- o^;-e arwles of the modified IiACA
(
four - digit - .series saat^=3 become nrogreasivel;,,- with incroasing airfoil thicla osa than those of t he convontional T:ACA four -digit-sories sect10Tis, a inre nro*io ncod forward mov=3-;unt of the wrokivIe:ic canter with lnci-ea.sinp; tliicL^-o p s would be oxoac ;3d for, t iesa airfoil sections; arA ti!,-) eoiaparative results for the FIAGA 0012-64 and 0010-34 sections teem to show this L-ond.
ItACA RM Ilo . L7122 13 COTICLUSIO:E _BLLic
Based u p on a two-climnsional invrstiL p ation of the aerokM
-it-serial airfoil sections characteristics of rive modified IInCA four-cIlLa
at Reynolds ciumbers from 3.0 x 10^) to 9.0 x 10` x , tho following
be drawn: cornclusio:rs ra y
y -in= lift charactari sties of the airfoil sections
1. The ma havia;,, normal-size leading-edge radii arA a. mxi.z= thic'alese of
r closely approxImt_-(I
12 percent chord located at 40 -norcent chord. ver,, those of NACA 64-sodas low-d.ra,,--^ , 3octiona oa correjpondinE th.'ieiales ; and carObar. .
2. `i'he Tram lift coefficicntua of Lho 10-:-;ercent-thick airfoils with ow-o artar rnorud-lie loadi.n^*-edgo radii euid' thiclMoss loca.-Wd. at 40 anti 50 portent chord were about 35 - nercont lower than thoso of aiooth MCA 64- series sections of corre3pordinc thiclsess
and cazdbar. For airfoils equipped. with ?0- ;ercent-chord split fla_.s
def lotted 600 , the ma:xima.m lift of the airf oils with om -quarter normal.-
size leadinc o-ed;e radii more warly approached that of PIA.CA 64-series
. Ro
airfoila a ughmss had no approciable offoct upon the nazi*_arm lift
of those airfoil-s.
3 . The r ni mom 1'•,OXI drag coefficients of the sirl o' is with MUM
thickioss at 40 percent chord_ and normal-size leadi^_-o3^ radii were
lusher than those of the corros-aond.inc I'MCA 64-:aeries sections.
Re ucins the le a s lnC^,-ed r )2i radius to ono- cuarter normal size and
movIns the position of =a.°.imum thic:mous to 40 and 70 percent chord
caused the Tji ,.i.m dray coeffici tc ba reduced to values about
ents the sama as those for corros-aoxuiin.e. FTACA 64- emd 66-3eries aoctions, rospectively.
RACA FV Rio . L (122
an^^;l©
4. Incroasas in the trailing-©dee p resulting from rear rd
mum thiciaiass causod sharp docroazes movemnt ci' the position of Lz_xLm in the lift-curve slope and proucunced forward novewnts of the aorodvnau to ceut.;r..
Langley Dlo=rim A©rorautical Laboratci-y National Advisory Committee for Aeronautics Ianr,ley Field, Va .
Lauz'enoo K. Lof tin, Jr.
Aeronautical ^aElneor nannoth S. Cohen Aaronautical irx;er Approved, Clinton H. Dearborn ChioZ of Fall calo Resea rch Division CJ3 NACA IN Uo. L7122 1. S tack, John, an i von Doenhoff , Albert E.: N at s of 16 Re late d Airfoil Soctionz at highSpoada . NAM Bop. Teo . 492, 1934.
Jacobs, Eastman N., Pins-arton, Robert M., and Graenberr, !Ia= : 2.
Masts of Relatad For mma-Caubar Airf cils in the Variable- Dvntait;^ ^lizar] `fu^nol. TVICA Re-p. No. 61.0, 1937.
3. -Loftin, Laurance K. Jr:: Theaorotical ani P.,-xari:iantal. Data for 110. 1308, a lhmibor of K. -kCA 6A- Series Airfoil Sections. M.CA `M 1947.
Abbott, Ira H., von Doenhoff, Albert E., and. Stivari, Loma S., Jr.: 4.
Sum m — of Airfoil Data. NAGA ACE No. L5C45, 1345.
5. `Z'haodoreen, Theodore: Theory of firing Sections of Arbitrary Shape .
NACA Rap. ?Io . 411, 1931.
and Pinkarten, Robert Ai .
6. Jacobs, East man. N . , Ward, Kienne th E - Y lue Characteristics Of 78 Related Airfoil Sections from Testa in the Variable -Density Wind Timm l. 14ACA Rap. No. 460, 1933 T. Langley 7 . von Doenhiol f , Albert E . , and " Abbott- ., I rank J. Jr.: TL_ , - Nwbulence Pressure aX=al. NACA M I g Two-Diimnsional Low-. o. 1283, 1947.
8. Schlichtink!, H., and Ulrich, A.: Sur Poruchnutae doe Umschle:Lles ln. r uiriar turbulont. Jahrb. 1942 der dautschen. Luftiahrtforsehung,
R. Oldonbaurg (Munich), pp. 1 8 -
1 35.
9. Purser, Paul E., and McKea, John W.. Wiz r id -Tunral Inwastioation of a Plain Aileron with Thickened and Bevelod Trailing EdCos on a 'Taoarod a y. -Drag Wing. PIACA ACn, Jati. 1943.
10. Jonas, Robert ., and Ames, Milton E ., Jr.: Wind-Tu=el Investi- 1.tion of Control-Surface Characteristics. V - T ie Uoo of a p
Bevoled ' linC EdC p
Lrai to Reduce the hi, p 4awrit of a Control Ow f aces . WICA A1 ; 2, March 1942.
11. Purser, Paul E. and Johnson, Harold S.: P facts of Trailing.-Edoa ., Modificationa on PitchinC. -Moment Characteristics of Aietoils.
11ACA CE 11o. L4130, 1944 .
NACA RM No. L7I22 e• TABLE I ORDP_v T ATES OF NACA AIRFOIL SECTIONS TESTED •s • I% •• NACA 0010-34 NACA 0010-34 a - 0.8 (modified), cLi - 0.2 [Stations and ordinates given in Stations and ordinates given in percent of airfoil chord] percent of airfoil chord] Upper Lower Surface Upper Surface Lower Surface Surface Station Ordinate Station Ordinate Station Ordinate Station Ordinate 0 0 0 0 0 0 0 0 I 68 .790 81 -.632 1.25 1.17 1.062 1.32 -.820 1.25 -.a44 .744 1.400 -1.400 2.5 2.5 2.407 1.6o8 2.593 -1.186 5.0 2.078 5.0 -2.078 4. 87 I 2.436 5.113 -1.714 3.0l^ 10 .5 - 3.oL^ 10 .5 7. 75 8 3 3 3.637 2. 4 12 lo.i^2 - }}55 6 4.523 2 8 6 15.1 -2.961 1 5 744 1 5 - 744 20 20 19.3 5.172 20.114 -3.312 :244 -, 244 N 29.917 5.980 o.C83 -3.684 39.95 9 0.045 4.833 4 5 .833 0 6.2 -3.721 9.99 66.186 5o.006 -3.526 60 60 4.433 4.433 70 . 061 69.93 6 1.915 2.1 -3.733 I ^p _2,767 80 63 2.7 77 80.100 3.700 -1.330 779.900 90 1.556 90 1.556 9o.076 2.044 8 9 -1.0 64 0 4 2 1.100 -.610 95 •856 95 -•856 94 . 95 0 95.
100.000 -.100 100 .100 100 -.100 100.000 .100 0.272 L.E. radius: 0.272 L.E. radius: Slope of radius through L.E.: 0.095 -64 NA'JA 0012 a = 0.8 NACA 0012-64 (modified), c Li = 0.2 [stations and ordinates given in [Stations and ordinates given in percent of airfoil chord percent of airfoil chord Lower Surface Upper Surface Lower Surface Upper Surface Station Ordinate Station Ordinate Station 1 Ordinate IStation T Ordinate 0 0 0 0 0 0 0 0 1.10 7 1.928 1.39 3 - 1 .686 1.25 1.813 1.25 -1.813 2.336 2.659 2.664 -2.237 2.5 2.453 2.5 -2.4 2.901.
3.267 4.823 623 - 5.177 1 -x•86 3
.5 7:5 7
7 7 z811^ 10 1} 2J^ 4.892 17:825
-4.8 15 10 .ibi -(.W
4:1
20 5.293 19.858 6.2 1 20.142 i _4.368 20 -5.293 5.827 0• -5.827 0 0 6.000 o - 6.000 7.279 -4.721 2 9.946 0.054 ^ 7.157 5 0 . 00 7 -4.4978 9 . 993 5.8 20 J•820 0.038 6.622 59.962 -4.o1 6o 60 70.077 55.662 69. 9 23 -3.296 80.120 1 1.253 9.880 i 3.320 -3.320 80 80 -2.383 90 1.867 90 -1.867 9o.091 -1.375 2.355 99.909 -.781 1.271 -1.027 1.027 95 95. 0 5 0 9+•95 0 -.120 .120 , 100 -.120 100.000 .120 100.000 L.E. radius. 1.582 rL E- radius: 1.582 Slope of radius through L.E.: 0.095 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS L_.
NACA RM No. L7122 TABLE I - Concluded • o ORDINATES OF NACA AIRFOIL SECTIONS TESTED NACA 0010-35 NACA 0012 [Stations and ordinates given in [Stations and ordinates given in percent of airfoil chord] percent of airfoil chord Lower Surface Upper Surface Lower Surface Upper Surface Station Ordinate Station Ordinate Station Ordinate Station Ordinate F0 0 0 0 --- 1.25 1.894 1.25 -1.894 2,5 2.615 2.5 -2.615 1.25 -.378 1.25 .878 2.55 2.5 1.2 67 -1.2 67 555 5 .0 555 5•0 - . 200 200 .
5.0 1.31,1 -1.81,1 5.0 68 05 3 1 63 -4.6 7.5 2.289 -2.289 1 0 5 4.5 7.5 10 2.667 to -2.667 9 20 -5.730 20 5.738 15 3.289 15 -3.289 -1 0 041 2 0 -6.002 2 0 6.002 4.788 3o 3 o - ^0 o -5.80 3 t55^.803 4.78 4.389 4 60 -4.339 boo -2.664 3.500 8o -3.500 80 2.6643 -2.100 0 1.
90 2.100 90 9 07 -.$ 1 , 17 8 -1.178 8 9 .8 95 00 -.100 100 ; -:182 126 100 100 .100 100 I 09272 1.58 L.E. radius: L.E. radius: NACA 2408 NACA 2110 [Stations and ordinates given in [Stations and ordinates given in percent of airfoil chord] percent of airfoil chord] Upper Surface Lower Surface Upper Surface Lower Surface Station Ordinate Ordinate Station Ordinate Station Ordinate Station 0 0 0 1.402 -1.448 1.098 1.694 1.128 1.380 1.372 -1.134 1.927 2.297 2.411 2.703 2.663 1.977 1.'493 2 .337 4.7 4 420 5.2 8 -2.1,82 2 3 -1.891 2.829 5.206 4.794 2, 09 4.169 7.733 1 2.11 - 7.217 7.'R .91 7.727 1 -3.016 7 66 10.290 3 3 7 4.
9-7 10 b6 15.278 -3.227 15 2.338 .222 - 14.7722 1.77 8 8 N7b 1 09 20.191 -2.320 5.320 6.666 25.1 6 I -3.236 I . 2.8]4 .852 5.677
2 2 5 .11 8 -2.239
-3.125 29.375 6.875 30.125 29.900 5.875 -2.125 3- 100 6.831 7 40.000 -2.33 -1.869 40.000 40.000 5.869 40.000 -2.46 50 .09 6.3 b 5 -1.535 49 9,95 1 50.0399 49.961 I L73 -z,o 55 5 5 6o.o68 4.820 -1.264 59.93 2 !
1 7 7 0.102 I 69 .93 -1.751 6 4.551 70.031 -•942 3.942 69.9 19 -1.074 80.097 60.078 a 2.858 79.9 2^2 -.636 779.9 0 3 '1:28'166 8 l 89.946 90.067 9 933 -•594 90.054 1.575 -.317 9 4 0 1 100.000 -.105 .105 -.084 j 100.0 100_000 ; .W , 10 0 .00 0 1.10 L i., Z. radius: radius: 0.70 I L.E. I L.E.: 0.1 ((Slope of radius through Slope of radius through L.E.: 0.1 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA RM No. L7I22 1.6 1.2 2 ]Ll .8 v ^ .t^ 0L 0 .2 .4 .6 .8 1.0 X/c X Ava/v (v/v) 2 v/v c c (percent (percent 4.8399 0 0 0 0 1.25 .756 .917 .958 1.36 2.5 1.120 1.023 1.011 .9 1.092 l.oi^5 .691 5.0 1.662 7.5 2.089 1.1 7 1.066 1.073 10 2.436 1.12 85 2.996 1.188 1.090 :387 1.098 326 20 3. 96 1.206 0 3.867 1.217 1.10 . ? j ^8 1.202 1.09 .19 ^o 4.000 550 3.884 1.185 1.089 .15 60 3.5L^7 1.16 3 1.079 .12- 2.987 1.127 l.0 2 .100 b ^ o 2.213 1.067 1 . 0 33 •074 1.6 •932 •965 .031
95 4
100 .0,30 0 0 0 L.E. radius; 0.174 percent c NACA 0008-34 basic thickness form.
Figure 1.- NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA RM No. L7122 °Kopf-I 1.6 G 3 NACA 0010 - 1.2 \\ v_ .8 v .4 o^ .2 .6 .8 1.0 x/c Ova (v/v)2 (percent c percent c V^ 0 0 0 0 2.321} 1.25 1.511 1.108 1.05 1.236 2.5 2.7L:4 1.245 1.11 .9 9 66 0 5.0 2-722 1.2 1.134 .69 3.178 1.277 1.130 556 10'5 3 1.2 9 1.127 . 533 .475 15 4.056 1.261
1.123 .37 7
4.111 1.248 1.117 .31b 3o 4. 8856 1.241 1.110 .241 5.000 1.242 1.115 •193 4.856 1.231 1.110 15^j 6 o 1.211 0 4.433 1.101 .126 0 1.1 4 5 1.074 .098 3.733 R 0 2.7677 1.09 1.043 .072 90 1.556 .980 .990 .045 .856 95 .912 •955 •030 .100 0- 0 0 L.E.
radius: 1.10 percent c Figure 2.- NACA 0010-64 basic thiclm ess form.
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA RM No. L7I22 •R • 1.6 1.2 2 .8 (V .4 0 .2 .4 .6 .8 1.0 X/C X (v/ V) 2 v, ^V Ova/V (percent c percent c 0 0 0 0 2.584 1.25 1.467 1.14o 1. 063 1.295 2- 5 1.967 1.273 1.123 0 5.0 2.589 1.271 1.127 4 7.5 2.989 1.252 1.119 .5551 10 3.300 1.236 1.112 .170 15 .7^6 1.213 1.101 .372 20 .0^9 1.200 1.09P .312 ^0 1.196
4. 3 1.091
j .239 40 1.212 4.19 1.101 .19 0 5.000 1.229 .15 1.109 ^ 0 4.867 1.234 1.111 .128 - H 4.339 1.226 1.107 .103 3.500 1.173 1. o8 .076 90 2.100 1.049 1.02 .046 .029 -95 1.178 •915 •957 100 .100 0 0 0 radius; L.E. 1.10 percent c Figure 3.- NACk 0010-65 basic thickness form.
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
^1_
NACA _ RM No. L7I22 . 1.6 ..
1.2 .8 .4 0L .2 .6 .8 1.0 x/C x Ava/V (v/V ) 2 v/V p ercent c ercent c 0 0 0 2.4 4 1.25 1.489 1. 1 0 1.06 1.29 2.5 2.011 1.2 6 1.11 5 5.0 2.656 1.2 6 1.134 6 9B : 3.089 1.282 1.132 .554 7.5 1.2 58 1.122 471 3.400 15 56 1.225 .372 1.107 20 .178 1.209 1.100 .310 0 78 1.189 .236 4.
1 - 0 20 0 4. 22 1.0 j 5 1 1.1 79 .190 ^ 5 1.184 1.089 .153 4. 956 0 6 0 1.211 1.102 . 12 5-000 .899 1.265 1 .125 .12 z 0 00 1.27 1.130 .080 4.
2.833 1.135 1.06 0 9 0 95 1.656 30 .960 .980 100 .100 0 0 0 radius; L.E. 1.10 percent c Figure 4.- NACA 0010-66 basic thickness form.
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
NACA RM No. L7I22
1.6
1.2
.4
I
.2
.4 .6 .8 1.0
X/c - x v/V Y -
QvA/v
(v/V)
0 0 0
3.857 1.25 . 44 .892 1.282 .944
2.5 1. ,00
1.011 1.005 .90
5.,0 2.078 1.113 .6. 8 1-055 2. 611
1.040 l.lb7 :581
10.5 3.044 1 1.095 .200 4 b 15 •744 1.238
1.113 .339
.244 1.256 1.121 .327
30 4.333 1.265 1.124 .249
5.000
1.253 1.119 .197 jo 4.856
1.237 1.111 .159
4.433 1.205 1.098 .127 1.17 1.076 .loo p 3.733
0 2.757
1.09 1.041 .073 90 1.35 •990 99 •045
95 . 910 .030
.954 loo .loo 0 0 0 L.E. radius; 0.272 percent c Figure 5.- NACA 0010 -34 basic thickness form.
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
k
NACA RM No. L7I22 1.6 1.2 vl .8 VIl
. 4
0L- 0 .2 .6 .8 1.0 .4 a/c va/v v/V (v/V )^ (percent c 0 0 0 o63 4.
1.309 1.25 .S7B .954 .977 2.5 1.267 1.032 1.01b .952
5.0 1. 44 1.043
1-03 7 •679 7.5 2.2 9 1.122
1.0
59 •555 10 2.667 1.141 1.068 .46 15 3.289 1.172 1.083 .3^2 20 .78Q 1.194 1.093 .323 0. .1 7
6 1.21) F 1.102
. 24 3 0 4.73 1.229 1.109 19 8
0 5 1.111 .1b2 000 1.2 2 0 .367 1.20 1.111 .ljl 1.227 1.10 .10L 70 4.389 .076 t30 3.500 1.176 1.084 90 .043 2.100 1.046 1.023 1.173 .920 .030 95 .959 100 .100 0 0.272 percent c L.E. radius: M Figure 6.- NA CA'0010-35 basic thickness form.
• NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA RM No. L7I22 . .
.
..
1.6 1.2 I
.4
0L 0 .2 , 6 .8 1.0 x/c Lva/V (v/V ) 2 v/v r x ^ 0 0 0 0 3.154 1.25 1.1 3 .865 .930 1.29951 5.0 1.122 1.0 9 9 633 2.49 .
3.133 1.136 1.o8 6o 7 .5 to .653 1.29 1.109 ^ 1. _32 1. 13^ ^:9 3 - 3 8 ^ 20 5.093 1.310 1.15 •329 0 55.800 1.329 1.1553 .250 o b.000 1.311 1.145 .198 20 5.827 1.231 1.13 .158 0 5 .
20 1.219 1.119 .128 7o 1^.^80 1.192 1.092 .098 1.9 2 1 - 0 5 1 90 :867 .029 95 1.027 92+ 94 .120 0 0 0 L.E. radius: 0 .391 percent c Figure 7.- NACA 0012 -34 basic thickness form.
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA RM No. L7I22 1.6 1.2 v • •4 0 :2 .4 1.0 .6 .8 x/c x Lva /d (v/V ) 2 v/V ( p ercent c) (percent c) 0 2.019 0 0 0 1.25 1.813 1.072 1.035 1.236 2.5 2.453
1.270 1.127 .9 5 2
5.0 3.267 1.33 0 1.153 .63 5
.813 1.325 1.151
10.5 1.322 1.1j 5 0 0
.L74
15 4t 1.313 1 il^6 .372
' 20 5.293 1.303 1.. 141 .315 0 5.827 1.297 1.139 .241 o 6.000 1.300 1.140 .19 ^ y 280 1.131 :154 60 5320 1; 1.115 126
o
^to 1.1' 9 1.090 .096
3.20 1.102 1.050 .070 90 1.367 6 •9 .993 1.027 .889 . 044 02 8 •93 100 .120 0 0 0 L..E. radius: 1.582 Percent c
Figure 8.-
NACA 0012-64 basic thicl^ajess form.
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NACA RM No. L7I22 •
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Figure Data for NACA mean a = 0.8 (modified).
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! I ..._. j -.08 - e 3.0 8° ?4I5 0.20c simulated s lit - - - - - flap deflected 60° .^ I _- "- -- - - - - NATIONAL ADVISORY x l vb 06 - - - COMMITTEE _ .'.._ FOR AERONAUTICS r c 1 nr—fTTYk 11.- Aerodynamic oharacteristios of the NACA 0010-34, a = 0,8 (modified), c ti = 0.2 airfoil section, 24 -inch chord.
F17gure I i 0.20c simulated spilt- i flap deflected 600 v 6 x lob '- V p Standardroughness i 6 x 10 6 - I I ._.^. c, 9.0 L_ x 106 6.0 3.0 Standard roughness boxlo6 m O x ..: - ... :.. _. f . .... ....:.:: -. r c I' R I lob _ J I v. . ..
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I 1 Aerodynamic characteristics of the NACA 0010-35 airfoil section, 24-inch chord.
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_ l mi a or ttc_ d n 24 -inch choxdA Figure 13.- Aerodynamic characteristics of the NACA 0012-64 airfoil section, • ••• • • • • • • •••• •••• •• •• ••• • t 0.20c simulated split flap deflected 600 R ,76 x 106 /Standard roughness N 6 x lob O r H .. .. ....
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6.0 x 106 L^ .0 Figure 16.- Aerodynamic chzracteristics of the NACA 2408 airfoil section, 4-inch chord.
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C) r Figure 17.- Aerodynamic characteristics of the NACA 2410 airfoil section, 24—inch chord.
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modified NACA 4-digit series CO NACA 4-digit series o0010-3 5 (present investigation; L0010- 3 - 0.2 O U010-34, a = 0.8 (modified), cli >00-series a 24-series 0 0012- 61t a = 0.8 (modified), c Zi = 0.2 00012 -61^, .012
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C^ l^ o .010 ft N .008 a^ NACA 4-digit series reference 4) _ ---- .006 NACA 64 -series r, (reference 4) .004 I -- U NACA 66-aeries m (reference 4) .002 Smooth NATfONAL ADVISORY - - -- - ROU Y1 COMMITTEE FOR AERONAUTICS 0 2 4 14 22 6 8 10 12 16 18 20 Airfoil thickness ratio, p ercent chord - Figure 1S.- Minimum section drag coefficients of several modified ITACA four - digit series airfoil sections, both with and without standard roughness, as compared with those of a number of NACA 6-series and NACA four-digit-series airfoil sections. R = 6.0 x 106.
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Modified NACA 4-digit series C^ 00010-3 NACA 4-digit series 00010- 3 (present investigation) Zi = 0.2 >0010-34, a = 0.8 (modified), c D 00-series 0012-64 <1 21{.-series '70012-64, a = 0.8 (modified), c Zi = 0.2 .14 ?lagged symbols indicate results with
r
standard leading-edge roug}uiess -Smooth 1- o 64- (reference 4) NACA series ----Rough e d ^ .12 a) a co a^ .1C U r - - 4-{ 4-digit series ( reference 4) NACA 0 .0E U NATIONAL ADVISORY W rn COMMITTEE FOR AERONAUTICS .066 22 2 10 12 14 16 18 20 Airfoil thickness ratio, percent chord Figure 19.- Section lift-curve slopes of several modified NACA four-digit-series airfoil sections both with and without standard roughness, as compared with those o^ a number of NACA ^4-series and NACA four-digit-series airfoil sections. R - 6.0 x 10 .
NACA RM No. L7I22 Modified NACA 4-digit series N-CA 4-digit series 0010-3 (present investigation) 0010-3 c 0010-34, a = 0.8 (modified), cLi = 0.2 00-series 0 0012-64 G 24- series V 0012-64, a = 0.8 (modified), cL i = 0.2 .ACA 64-series (reference 4) 2.8 i Smooth cLi .2
Izxz
ci o 2.4 U ^D hough O CLi U 4.
1~ ^ .2 U 2.0 ri U q _ •U NACA 64- series W 1.6 (reference 4) W a) i S:,00th U cLi l.c .2 Eough c U a^ cLi ^ I P ^- ID .2 Snooth -Rough
T7_
Z FlaEged symbols indicate results with standard leading-Pds;e roughness NATIONAL ADVISORY M FOR AERONAUTICS -COMMITTEE -COMMITTEE I 20 24 8 12 0 4 "irfoil thicla:esz7 ratio, percent chord Fivure 20.- "laximum section lift coefficients of several modified ',ACA four-digit-series airfoil sections, ooth with and without standard ro • ^._Yness and S lit flaps, as k k compared with those of a number of NAC_ 6l;- series and :::.C;-. four-di€it-series airfoil sections. F. = 6.0 X 106.
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Modif ied NA,01k -it series
4-di
y
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U G •24 i G series (reference ^) NACA 4-digit i +-..22 •4 v ' I a, NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS y • L 16 20 24 0 0 4 8 12 Airfoil thickness ratio; percent chord Figure 21.- Chordwise position of the a.erodyna.mic center of sever.,,l modified NACA four-digit-series airfoil sections as compared with those of a number of NACA 64- series and NAOA four-digit-series airfoil sections. R = 6.0 x 106.
NACA RM No. L7I22 INDEX Subject Number
Wing Sections - Profile Variables .2.1.2
1.2.1.2.2
Thickness - Wing Sections
Thickness Distribution - Wing Sections 1.2.1.2.3
Surface Conditions - Wing Sections 1.2.1.2.5
Profiles, Designated - Wing Sections 1.2.1.3
Flaps, Split - Wing Sections 1.2.1.x+:2
Reynolds Number Effects - Wing Sections 1.2.1.7
ABSTRACT Theoretical pressure distributions and measured lift, drag, and pitching moment characteristics at three values of Reynolds number are presented for a group of NACA four-digit-aeries airfoil sections modified for high-speed applications. The effectiveness of flaps applied to these airfoils and the effect of standard leading-edge roughness were also investigated at one value of Reynolds number.
Results are also presented of tests of three conventional NACA four-digit-series airfoil sections.