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1
NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
TECHNICAL NOTE
No. 1278
LOW-SPEED TESTS OF FIVE NACA 66-SERIES AIRFOILS
HAVING MEAN LINES DESIGNED TO GIVE
HIGH CRITICAL MACH NUMBERS
By Albert E. von Doerihoft, Louis S. Stivers, Jr.,and James M. O ‘Connor
Langley Memorial Aeronautical LaboratoryLangley Field, Va.
Washington
May 1947
——
.
TECH LIBRARY KAFB, NM -
llllllllllillllllllllllllllUllllllEllw45nl e
.—
NATIONAL ADVISORY UCWWITTEE FOR AERONAUTICS
TEC~ICAL NOTE K). 3276
LOW-SPEED TESTS OF FIVE
HAVING MEAN LINES
HIGH CRITXCAL
NACA 66-S3RIES AIRFOILS
DESIG12ZDTO GIVE
}JACHNUM~RS—
By Albert E, von Doenhoff, Louis SYS1.Yi~, Jr. ,aridJames M. O!Connor ............
SUMMARY
The possibility of developing an _Qirfoil to carrylift without decreasing the critical Mach number belo~’fthat of the basic thickness form at zero lift has beeninvestigated, Low-speed tests of five NACA 66-seriesairfoil sections b.a~i~ a thickness-chord ratio of 0.16were mqde in the Langley two-diw.ensional low-turbulencepressure tunnel, w designing the mean line to carryload over the portions of the airfoil section having lowinduced velocities an effective design lift coefficient(the lift coefficient corresponding to the center ofthe range of high critical Mach numbers as obtainedfrom the experimental pressure distribution) of approxi-mately 0.1 was obtained for several airfoil sectionswithout causing the maximum predicted critical Machnumbers to be appreciably less than the critical Machnumber for the basic thickness form at zero lift. Themaximum lift coeff’ici.entsend the drag coefficients inthe low-drag range were ap roximately the ssme for tlnese
%airfoils as for the NACA 6 -series airfoil sectionshaving the same thickness and a>proxirnately the sameeffective design lift coefficient with the uniform-loadmean linec The low-drag range at a R~ltisnumber. .-of 9 x 106 decreased with increase in-~=ign lift coef-ficient above a value of 0.2. The pitching-momentcoefficients were larger than those of airfoils havingthe same effective design lift coefficients with the
..___
uniform-lead mean line but were not ne”tiilyso large asthose corresponding to the design load distribution.
Reco~endations concerning the use of the airfoilsat high speeds cannot be made because of the lack oftkst data at high filachnumbers.
L
2
IW!TRCIDUCTION
NACA TN NO. 1276 8
I
A large amount Q2 work has been done on the problemof designing eirfoil that nave high critical hlachnumbers. zThe NACA 1 -series girfoi.ls,presented inreference 1, have a ttikckness distribution that givesunusually high cri$$~al Mach numbers for a giventhickness-chord ratio. This series of airfoil sections,however, has high crttica~ ],lachnumbers over only alimited,range of llft ~cjef~icients. Many of theNACA 6-series airfoils, data for which are presentedin reference 2, have critical Vach numbers somewhatlower than those for the corresponding airfoils of theNACA 16-series but have high critical Mach numbers overa considerably larger range of lift coefficients. Theforward portion of the NACA 6-sertes sectfun-s are
.-
designed so’that the pressure distribution forward ofthe point of minimum pressure becomes esstifitiallyflatat the extremities “ofthe ran~e cd’lift coaf$’i.cientsfor low drag. For .a,give.n.positi~n of minimum pressureon the basic thickness form, this design cnndition givesan airfoil.shape that has a minimxm increase in nlaximum- ?.velocity ratio throughout the range of lift coefficientsfor low drag.
●
Because the mean line corresponding to a uniformchordwise distribution of load at the design lift coef-ficient (a = 1.0) has the highest possible criticalspeed for a given lift coefficient, this mean line hasbeen most frequently used es the mean line.ofNACA ~6-ser~es and KACA 6-series sections. Although theuniform-load mean line has the optimum critical-speedcharacteristics for the mesn line itself, an airfoil offinite thickness having this mean line wi~l not necessarilyhave the highest pos~ibls critical speed for a giventhickness and des5.gnlift coefficient. ThQ criticalspeed of an airfoil section is determined by ths maximumvelocity oc~urring on the airfoil surface. -.If the meanline for a given thickness distribution is designed soas to cause the airfoil to carry lift--over the portiQnsof the chord where the velocity is less than themaximum, the airfoil will then be able to carry somelift and not have a velocity ratio greater.than themaximum for the basic thickness form, In @der for theairfoil to carry the largest”amount of liftiwi.thout
—.
decreasing the critical Mach number below t“hatfor the
.NJICAT’l? i;os 1276 3
=Y*
basic thiclcnessformat zero lift, the pressures overthe upper surf’aceof the air~oil uust be uniform andthe pressure coef~icie:ltmust equal that Of the basic.thicluloss form.
For the portion of an NACA 6-series airfcil for-ward of the position of minimum ~i’essure, the loaddistribution associated with angle of attack is optimw:lat the lift coefficient corresponding to the upperextremity of the lov~-drag range beta.use”at this liftcoefficient the.pressure distribution ~vcr the.upper ~surface becomes uniform fro-mthe leadin,q ed~e back tothe ori~inal ptisition of minimum pressure. w order notto disturb the pressure distribution ovor the forwardportion of the aii’f’oil,the optiznzmmean line for highcl’itical lilach numbers should be deslbned to give zeroload fron the leading ed~e to the po+intof minimumpressure on the basic thickness for~,~and to give a loaddistribution correspondfig to uni~orm”pressure over theupper surface from this point to the trailin: edge.Such a load distribution ordinarily corresponds to alarge flilite load at thetrailti: edge. Previous.experience wi-:h airfoils havinS mean lines desi~efl togive finite loads at the trailing edge indicates thatthe load’distribution over most of?the chord 1s su3-.astantially as specii’iedbut that the finite load At thetrailin~ e~ge is not realized in practice.
The purposes of the present investigation are (1) todetermine experimentally the extent to which the methodsjust described are effective in increasin: the desi~nli~t Coeti’icignt o.fm airfoil witi:loutdecreasing thecritical Mach numbgr apprecia-oly below that of the basicthickness term and (2) to determine the effects of thecorresponding LKJW3Ud “tymeof load .distrib~tion oncharacteristics of the a~rf’oil sectic~ other th= thocritical Mach number, such as ‘:pitchinzmoment, maximumlift, and drag. The present investigation includes low-speed tests in the Lcmgley two-dimensional Iow-turb-dmacepressure tunnel of five airfoil sections having theNACA 66(215)-016 basic thickness form. Three of thesesections have mean lines designed to carry variousamounts of load bac’k of tl~eposition of minimum pressure.of the basic thickness form (0.6 chord) at the desi:n”lift coefficient. The other tvo air:oils have mean
* lines desi~ned to carry a part of the total loadUniformly over th-eentire chord. Tiietests consistedof measurements of lift, dra~, and pitchi~’ moment at
.-
. ..—
4 NACA TN No, 1276
Reynolds numbers of 3 x 106, 6 x 106, and 9 x 106and at Mach numbers less than 0.17. Low-speed pressuredistributions f’oreach of the airfoils were determined
at a Reynolds number of 6 x 106 through a range ofangles of attack corresponding to a range of lift coef–ficients from large negative values to values beyondthe positive stall..
Definite recommendations concerning the use of’theairfoils at high speeds cannot be made because of thelack of test data at high Mach numbers. Additional dataare also needed on’the application of lateral-controland high--lift devices because of the unusual shape ofthe airfoils near the trailing edge.
a
c
cd
Cdmin
Ct
clmax
c1i
Cma.c.
% c/4
Ho
Mcl?
sYMBoLs
mean-line designation, fraction of chordfrom leading edge over which design loadis uniform
chord
section drag coefficient
minimum section drag coefficient
section lift coefficient—
maximum section lift coefficient
design section lift coefficient _
moment coefficient about aerodynamic center
moment coefficient ‘aboutquarter–cho;d point
free-stream total pressure
critical Mach number
.
9“
*
NACA TN I?o. 1276 5
P local static pressure
Clo free-stream dynamic pressure
R Reynolds number
(0 )H –Ps pressure coefficient
qo
x dtstance along chord line measured from leadingedge
Y distance perpendicular to chord line measuredfrom chord line
a. section angle Gf attack
DESIGN OF TllXAIRFOIL SECTIONS
The basic thickness form for all the airfoils testedin this investigation was tineNACA 66{215)–016 airfoilsection, which has minimum pressu~e at c).6c from theleading edge and a thickness--chord ratio of 0.16.. ASpreviously dtscussed, mean lines were desired that have~eyo load from X,/c= o to X/C = 0,6 and linear].yincreasing load from this point to t@e trailing edge,Because the relations obtained from tinetheorj of thinwj.ng sections are linear, the theoretical mean linesand.load distributions can be obtained simply byaddition of’the ordinates and the correspondlfi.gvelocityincrements of component mean lines.
The desired type of load distribution is obtainedby a combination of a uniform-load mean line (a = 1.0)with a mean line havin,g uniform load from the leadingedge to 0.6c and linearly decreasing load from 0.6c tothe trailing edge (a =:0.6), In order for the loadto be zero over the forward portion of the airfoil, thedesign lift coefficient for the mean line of the type
. a = 0.5 must be -0.8 times the design lift coefficientfor the mean line of the type a = 1.0; in order for anairfoil to have a design lift coefficient of 0.2”, the
4 sum of the design lift coefficients cf the component .mean lines’must.equal this value. These conditions
6 NACA TN No. 1?76
are satisfied b a mean line having the f’ollowincomponents: (l~a mean line of the ty e a= O.
?)E With
a des~.gnlift coefficient of’-0.8 and 2 a mean lineof the type a = 1,0 with a design lift caef’fici.entC?f1.0. ‘fhedesignation of the airtoil having thechosen basic thickness form and the aforementioned meanline is as follows: .
ra= 0.6, Ct = +,8NACA 66 (215)–216
1
i- Ia= 1.0, Cli = 1.0
J
Further details of the numbering system fop this typeof des~gnation are discussed in reference 2, -In orderto determine the effects of varj.ation in camber, twoadditional airfoil sect~.onshaving theoretical designlift coefficients of 0.3 and 0.4 ~-eroderived. Theseairfoil sections are: .
--- +a= 0.6, Cti = –1.2
NACA 66(215)–316
{
\a = 1.0, Czi = i.5 J
/
Ja= (.).6, Cz =–1.6 !
NACA 66(215)–4z6 i >
1a = 1.0, Cz = ~,o i
i. 1
For the three airfoils the theoretical pressuredistributions at the design lift coefficient-, presented ““”in figure 1, indicate that even the airfoil designed fora lif’tcoefficient of 0.2 has theoretically a slightlyhi.gaermaximum value of the pressure coefficient S,and hence a lower value of the critical Mach number, thanthat-for the basic thickness form. The experimentalpressure distributions, however, were not expected toshow this decrease in critical Mach number because offailure to realize fully the theoretical load dis–tribut~on.
Two mor~ airfoils were investigated for the purposeof determining the extent to which the critical Machnumber characteristics of airfoils cambered with auniform-load mean line could be improved by increasingthe portion of the load carried by the rearward part of
.
NACA TITNo..-*’
,.
.
the airfoilthe forward
=76 7
while holding uniform the load carried bypart* The ttioairfoils are:
laL
= 1.0, c2~ = 0.7’)
(EL= 0;6, c~i = -om3\NACA 66(215)*216 K
[ a = 1.0, cL~ = 0“5J
The contours and corresponding theoretical loaddistributions for the five airfoil sections c~nsideredin this investigation are given in figme 2. Ordinatesare given in tables I to V.
MODELS AND TESTS
,Models of the five airfoil sections were built of
mahogany laminated in the chordwise direction. Each
model had a chord of & inches and a span of 35& inches,3The models were prepared fol:standard tests in sheLangley two-dimensional lcw-turbulence pressure tunnel(TDT) in the manner described in reference 2.
Lift data were obtained from measurenen~s of thepressure reactions on the floor and tailing of thetunnel, drag data were obtained from measurements bythe wake-survey method, and pitching-momant data were -measured with a torque balance, Details of the methodsof obtaining the data we given in the appendix ofreference 2. -—-.
Lift, drag, and pitching-moment data were obtainedF , ,at Reyimlds numbers of 3“x 10b, 6 X 105, and 9 X 10bfcr rmdels in a smooth condition, pressure-distributiondata for each model were obtained at a number of anglesof attack corresponding to a ra~e of lift coefficientsfrom large negative values to values beyond rn&hinun lift;thesa data were obtained for the smooth models at a. .
Reynolds number of 6 X 106. with a standard roughn6ssapplied to the leading edges of the rnodals, lift and+-., ., ...-.
.
6drag data were”’obtained at a Reynold’s number of 6 x 10..This rcughness consisted of approximately ‘0.011-inchgrains of carborund.um applied 1+ the airfoil surfaceover a flurfacelength of.0.08c meas~ed fro-mthe leadinged~e on both upper and lower surfaces, The grains werethtnly spread to cover ~ to 10 percent of this area.
RESULTS
The experimental pressure distributions are~rosented in figuz-es 3 to 7. Lift} drag; ~d--pitching-Jmoment data for the five aiio$’uilsections are preaent6din fi~uros 8 to”12*” The force dat&”h~ve been correctedfop the constricting
[effects-of’the tunnel walls by
equations (37) to ( O) in the appendix of reference 20For the present airfoils these equations are reduced tothe following simql.ifiedforms, where.the primed valuesrepresent the values measured in the tunnel:
ao = 1.015CJO! ‘
Corresponding corrections havo been applied to thopressure-distribution datai
,
DISCUSSION
Becausa only low-speed data were obtatned~ corn-parisons of the v’arious””air?oilsections atiemade onthe basis 01 predicted critical Mach inimbers. Ingei~e.ral,critical Mach nuiibers (the critical Mach numberIs defined as that free-stream Mach number at which thelocal velocity of’sound is first attained) predicted ~from low-speed experimental pressure distributions arein ~ood agree~ent with htgh=apeed-test results, Thiscritical Hach humber is somewhat lower than and is tobe distinguished from Lhe Mach number corresponding bothe force breaks,
.
v“ I
.
*
.
NACA
were
TN NO. 1276 9—
The critical Mach numbers presented in figure 1?obtained from the theoretical and ex~erimental ow-
speed pressure-distribution data by the van Ke!rm<nmethod by use,of the curve presented on page s83 of’supplement IV, reference 2. The curves of the criticalMach nwnbers corresponding to the theoretical andexperimental low-speed pressure distributions of thepresent airfoil sections are cdnpa.red with sinilar datafor ‘airfotl sections having the sane basic thicknessform and the same effective design lift coefficient (thelift coefficient corresponding to the centetiof therange of high oriticel Mach numbers as obtained from the —
experimental’ pressure distribution), but having a uniform-load mean line (a = 1.0). Although the theoretical .wd -.effective design lift coefficients may be seen to differ “-considerably for the newer airfoils, tined~ta of refer-ence 2 show that the effective and theoretical designlift coefficients are substantially equal for the air-foil sections cambered with the uniform-load mean linefor moderate design lift coefficients. Numerous pressure-distribution measurements heve also shown that the theo-retical and experimental low-speed pressure distributionsfor the latter airfoils are in good agreement at low andmoderate lift coefficients. .
The data of figures 13(a) to (c) show that, for theairfoils having zero load from tineleading edge to theposition of minimum pressure, the effective design liftcoefficients are less than the theoretical coefficientsand this discrepancy “increases with increase in camber.For these airfoils the maximum critical Mach numbersobtained from the experimental pressure distributionsare equal to or greater-than the maximum critical Machnumbers obtained from the theoretical pressure dis-tributions. The critical Mach numbers obtained fromthe experimental pressure distributions increase withincreasing lift coefficient in most of the range ofhigh critical Mach numbers and are a maximum between” ““-’-- --lift coefficients of 0.2 and 0.3. These maximumcritical Mach.nunbers are approximately 0.0Z5 greaterat these lift coefficients ‘than the critical Machnumbers for the airfoils having the uniform.-load meanline. The results presented in f2gures 13(a) to (c)indicate that the airfoils have an effective design liftcoefficient of approximately 0,1 and have maximum -critical Mach numbers at normal lift coefficients forhigh speed that approach the maximum critical Mach
.
,,
.
10 NAdA TN NO. 1276 v
number ~or the basic thickness form at zero lift. Thedifference between .the critical Mach number .-ofthe basicthickness form and of the cambered airfoil sectionsincreases with increasing” camber,
For the”two airfoil “sections c:~rrylng a-portion of -the load from &lieleading edge to the position of minimumpressure (figs. 13(d) and (’e)),the predicted crltlcalMach numbers obtained from the exper-imental.and theo-retical low-speed pressure distributions azze-much closerin agreement than for the other three airfoils. Thesetwo airfoil sections show small gains in critical Machnumbers as compared with the airfoil section_shaving theuniform-load mean line.
.
A comparison of the theoretical pressure dis-tributions at the design lift coefficient i~iththeexperir.ental pressure distributions having a load dis-tribution over the forward portion most nearl
Jlike the
design load distribution is given in figure ~.Figures ~(a) to’ (c), which present data for tk.ethreeairfoils with zero load from x\c = O to x\c = 0.6, showsuccessively greater discrepancies between the experi-mental and the theoretical pressure distributions withincrease in design lift coefficient. At the designlift coefficient, theoretically, no I“oad should beobserved over the forward portion of these airfoilsections . This condition is not fulfilled begause ofthe failure to realize the design load over the rearwardportion of the airfoil. In figures U(a) and (b) thevalues of maximum pressure coefficient for the experi-mental preseure distributions are in.good agreement withthe theoretical values. The data presented infigures l&(d) and (e) indicate that the agreement betweenthe experimental and tneoretice,lpressure distributionsbecomes progressively better as the load on t~e rearwardportion of the airfoils is decreased:
-.
The low-speed aerodynamic data for the airfoils con-sidered in this investigation are given in I’igures 8 to 12and are summarized in table VI. A comparison ofithesedata with those for the NACA 66-series airfoil sections(reference 2) havin~ the same thickness and approximatelythe same effective design lift coefficient with theuniform-load mean line indicates no lar~e difference inthe maximum lift coefficients or in the”drsg
in,the low-drag range. At a Reynolds number
coefficients
Of 9 x 106
.
-a“
the “newer airfoils having .theoretical design lift coeff-icients greater than 0.2 show a prog~essive decrease .of the low-drag uange with increase illcamber and showa sornewhat larger initial increase in drag coefficientsat the upper end of t’herange: of.~OW drag coefficientsthan t~~ecorresponding NAGA 66-s~ries airfoils with theuniform-load mean line. The test data for the newerairfoils show a jog in the lift curves at the uppen endof the”low-drag,range. The magnitude of the jog.decreases as tihe’~ynolds .number.is increased from .“
63x”lotogx lo. For”the airfoils with a theoreticaldesign lift coefficient of 0.2 the jog in the lift curve
at a Reynolds number of 9 x 106-is Very small and has . “ ----approximately the ~~ie magnitude .as the jogsfo.und for theNACA 66-series, airfoils (reference Z) having approxi-mately the same ef~ective design lift coe~ficient and theuniform-load mean line.
—Tb.e jo&, b.owever) for ‘the new”~fi
airfcils with a theoretical-design lift coefficient 6of 0.3 and 0.4 is-greater at a Reyfio?-ds”ati-b~rOf 9 x 10than tinejogs for tinecortiespondikg l~A~A66-series air- - ‘foils (reference 2) with”the uniform-load me~~ line Orfor the other airfoils of the presentg~.per.. The magni-tude of the jog appears to ihcr’ease’with”increas~-in
—
theoretical design.lift CO.effiQie,ntfl . ,, , .: .__: ___ ~----
The pitchi~-mo~.ent coefficients” ‘of’the newer air-foil sections are larger than .thos’efor the NACA 66-seriesairfoils (reterence 2) having t’nesane effective designlift coefficient with the uniform-load mean line butare not nearly so large as the pitching-moment coef-ficients corresponding to the theoretical load dis-tribution. .. .
CONCLUS1ONS . :, _ : .“ ‘-
-~lthough, yecommendatLons concerning the,us~.of the“airfoils at.high-speeds .c@not be “rna~e,becav.se of,:thelacK of test ~ata at higk~llach tiumbers~’._loy-sp.e_e.d‘testsof five ITACA ~6-series airfoil .iect~ons having meanlines designed to give high critical Mach numbers indi-cated the following conclusions; .-
1. An effective design lift coefficient (the liftcoefficient corresponding to the center Of the r~ge of
12 NACA TN NO. ti76
high critical Mach numbers as obtained from the experi-mental pressure distribution) of approximately 0.1 wasobtained for several airfoil sections without ceuslngthe maximum predicted c~ltical Mach numbers to beappreciably less than the critical Mach number for thebasic thickness form at zero lift.
2. .The maximum lift coefficientsiand the dragcoefficients in the low-drag range were approximatelythe same for these airfotls as for the I?MA 66-seriesairfoil sections having the same thickness and approxi-mately the same eftective design lift coefficient withthe uniform-load mean line.
3* The low-drag range at.a Reynolds number of
9 x 106 decreased with increase-in deslgnllft coef’- ,ficlent above a value ot 0.2.
-.
J. The pitching-mo~nt coefficients were largerthan those Of airfof,ls having the same effective designlift coefficients with the uniform-load mean line butwere not nearly so large as tln~secorresponding to thetheoretical load distribut$onf
Langley Memorial Aeronautical Laboratory
1.
2.
—.National Advisory committee for Aeronautics
Langley Field, Vs., l?ovember 29, 1945
REFERENCES
Stack, John: Tests of Airfoils Designed to Delay
E::;Ksst::;izt%b& ~~:: :X:i ““
Abbott, Ira H., von Doenhoff, Albert E,, and Stivers,Louis S., Jr.: Summar of Airfoil Data.
tNACA ACR No. L5C05, 19 50
.
v
.
.
NACA TN I!Jo.1276 13.
‘rAsIE I
OSOINAISS OF ‘Et5
TAs18 11
0FDIi4A7SSOF =
‘ni = -1.2
}%i = 1.5{a = 0.6,
UAOA66(215)-216 * = ~oo,{
= 0.6,HACA66(215)-316a
a = 1.0,
U N7?OILW TIOU AIRFOILSWTIOll
[Stations and ordinatea given inperoent of alrfoll ohcmfj
CStathm ad C@Mnatea given inperoent of airfoil ohmdl
Uuuer aurfaoe Immr surface Uppm”Surface kmnr eurface
~rdhata
.
.
L.E. radlua: l.~jS1OPO of radius throu@ L.E,
L.E.radius;1.575Slopeof rad~us through L.%s -0.02(
NATIONALAovMSIS III
o~NAES OF ~
mm Iv COMMITTEE FM AERONAUT)=
0~llAm5 OF THS
{c = 0.6, % = -1.6
nAoA66(215)+16= = ~.o, ~zii }= 2.0 {
a = 0.6, %L x -&sifACA66(215@6
a = 1.0, ali= 0.7}UFFOIL 8EJ’IT.oN UmOIL mc moH
[Statlonm and ordfaatea given inperoent Or a~rron OhOrd)
[StatfOn8md e:peraent of
Upper surfaoe
mtea given bv?o11 chord]
km r surraae
.
.,,.
-
14
T.mLs v
OSOISAW OF TIE
{ }
a = 0.6, oq = -0.3SAOA 66(2151-216
6 = 1.0, Oq z o.~
AISFOILSX~OH
[Stations and ordlmtea given Innaraant of airsM.1 olmrd]..— _–.
Upper surfaoe-stat100 ?dlaate
Lower mrfaoe
)tatlon
-.—
NACA TIN No. 1276
.
.
NATIONAL ADVISORY~glrm.s Fm ASSOSAUTU
--—
-..
.
-.
.
#“
# , Al,
TAB~ VI
SWAIW OF AERODYNAMIC1- .7LR = 9 x lob]
NACA airfoil section
{
8 = 0.6, Cti = -0.866( 215 )-216
a = 1.0, CL: = 1.0 }
{
a = 0.6, Czi = -1.266(215)-316
a = 1.0, ox{ = 1*5}
“{66(215) -@6 a 1=0.6, fill = -1.6
a = 1.0, CL: = 2.0
66( 215 )-216
{
a = 0.6, c1
}
= -~”5ia = 1.0, CL = 0.7
.i
66(215)-216{
a =0.6, Ct = -0.3i
a= 1“0’ CL* = ~“5}
1R= 6X106.
cd
mln
0.0034
0.0031
0.002?7’
0.0035
0.0034
c Zm
Smooth
1*47
1.51
1*55
1“55
1*43-.
cmRoy a.c.
0.99 -0.050
0.96 -0.075
---1-1.01 -0.100
1.05 -0.052
NATIONM ADVISORY
COWTTEE FOR AERONAUTICS.
2!
Pcm
Fig. 1
s
NACA TN No. 1276
.
.Z
1.8 I I IWper surface
_- --- --
1.6 #- r/ F I – I \
1.4 — - — — — — . \/ ~ \ /
i\
/\
=1~= O.J$ <1.2
/
/?;
\
\\
\
l.o
/
ox = O.y1
\ /\ \/
\ \.8 --
\cl
i
\ ‘ \ /\
.6 \\
Baa10t~ckneasfoFmr \ \
.4 \\\ \\
I I ,Lo-rsurraoe <’ \
.2 \
I
\
$\NATIONAL ADVISORY \
COIIHITTEEFDRAERMAUTICS
00 10 20- 30 40 50 60 70 BO 9CI &Peroent ohord
F@we 1.- ~eoretloal pressure MS trlbutlons for the three airfoilseotion8 ~ving zer loadlng from O to 0.6 alrfoll ohord and forthe NACA 6t215)-ol~ baelC thiokness few.
I
NACA TN NO. 1276 F’@.2
*
.
Theoreticalohordwise loadings
.Airfoil cent ours
{= 0.6, OLL = -0.8
NAOA 66(2151-216 : = 1.0, Oti
= 1.0 1
{=0.6
MAOA 66(215)-316 : = 1-o: ::= -1.2
1= 1.5 1
{
= 0.6,Ozi= -1.6NAoA66(z15)-l@~sl.o, c~~= 2.0 1
-.
{
= 0.6,’11=-0.5NAOA 66(215)-=6 :s1.0, oli s 0.7
}
{= 0.6,OL1=-0.3
NACA 66(2151-216 := LO, c~i = 0.5)
NATIONAL ADVISORY
COWRTTEEFMAERONAUT KS
..—..
-.
—.
.
Figure 2.- theoretical ohordwiee loadlngs aud contours for the five airfoil nections.
.
.?2
Lo
1.6
1.6
Lb
1.2
a
1.0
.0
.6
&
a
0
z
1. I I I I I I
I 1 I I I I I I I I I I I IM
n Illllllllllllll”k, ,\ ,
..9 I PI
w! 1 1 1 I 1 I I 1 I I I I I 1\ )
.6I
SY I111, ,,,
iA
.2
x ‘* 4
. ●, f.. ,
.,
2.0
1.8
1,6
1.4
1.2
B
1.0
.0
.6
J.
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{
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airfoil section.
Figure l?I .- Comparison of the predicted critical Mach.numbersobtainedfrom the theoretical-and experimentallow-speedpreseuredistributions.
. ---
—
Fig. 13b INACA TIN INo.1276.
-..r!.
l’~:::~:!::l::”.[i’::]..I :..] [ .:.1 I:“:l:..J,::. .. . .,.4 . . ..+ . ...1..--1 .1! I I I 1 I 1 ,
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M
-.......-.:.. for the HACA.66(215)-116airfoil ,;:;:::;,,!::::.,,,...,............;, sec.tl.on
.—.... ..........,...,....,.::h —__ ~~nm +hnfivn+4nn7 nnmmIIva il~e+w+h??+~nn
:..- .,. . . .. .. . .... . ..
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{(b) NACA 66(215)-316~~’~”j’ ;;; ~-:”~
}airfoil6ection.
● , .Figure13.- Gontinuedl
NACA TN INo.1276 Fig.13C
,..,,.,,, ...,. . . ... ..,,.:., ...,. , From experimentalpressuredistribution.,,:...::....:::::...,.,..,.,.,..... . — — From theoreticalpressuredistribution:.:.;,.:,.:!-––-– Fran theoreticalpressuredistributionfor the lUICA66(215)-(1.3)16airfoil,.seotlon
—–-From theoreticalpressuredistributionfor the NACA 66(215)-016 airfoil
-..._
Fig.13d NACA TN No. 1276
.,,:..~.. r.~. ~ ,,, : , , -,,. :,,:... . . . . .
. . . . . . . . . . :-.:....::: !:: 1 I
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Figure13.- Continued.
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.—J I
NATIONAL ADVISORYCOH~lTTEEF~ AERONAUTICS
10 20 30PerOentahord .-
{ }= 0.6, ‘zL = -0.8
(a) NACA66(215)4Q6 ~ =LO, all = LO airfoil section.
I
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Figure 14.- Comparlson of the expertientalpremwre dlatrlbuti.on●t the effeotlvedesign lift amfflolent with the theoretloalpressure distributionat t@ dealgnlift coefficient.
●
NACA TN No. 1276Fig. 14b
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COAWTTEEFORASRONAUT4CS\
( )
O(j 10 20 30 40 50 w 70 60 90 -Jo
Peroent chord .
{ }(b) EM3A66(215+316 a = 0“6s ‘a%‘-;:;
a = Lo, Otl = alrfoll seotion.
FlgtH lk.- Ccmtinued.
Fig.1.4c NACA TN No. 1276
I
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NATIONAL ADVISORY \
COMMITTEEFM AERONAUTICS\.
Og 10 20 30 40 so 60 70 80 92 100Peroent ohord
{ }(o) NACA 66(215)@6 ~~~”~s ~~i ~ ‘$: airfoil section.
■ s 1.
Figure 14.- Continued,
NACA TN No. 1276 Fig. 14d
.
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1.6
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NATIONAL ADVISORY,. 2 COMMITTEEFORAERONAUTICS
@c1 10 20 30 40 50 60 70 80 90 100
Peroent chord
{ }
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%~ = 0.7
Flgwe 14.- Continwd.
Fig.14e NACA TN No. 1276
1.8 .
1.6
1.4
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50 60 70 80 90 l’boPeroentohord
{(e) HAGA66(225)-226a ‘0”6’ ~~t‘-0S5
a = 1.0, = 0.5}airfoilneotion.
1Figure4.- c~ol~edo
