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NATIONAL ADVISORY COMMITTEE

FOR AERONAUTICS

TECHNICAL NOTE 3244

AERODYNAMIC CHARJICTERISTICS OF THE

NACA 64-010 AND 0010-1.10 40/1.051 AIRFOIL SECTIONS

AT MACH NUMBERS FROM 0.30 TO 0.85 AND REYNOLDS

NUMBERS FROM 4.0 X 106 TO 8.0 X 106

By Laurence K. Lofh, Jr.

Langley Aeronautical LaboratoryLangley Field, Va

Washington

August 1954

.—. .,

I

—.--+ ___ .- ..<.. ..-~. . . . . . . . ,. .-.. . . .. . . . . . .. ---- ._. . . . . . . . . —>.- .—.

https://ntrs.nasa.gov/search.jsp?R=19930083926 2020-04-29T13:13:19+00:00Z

TECH LIBRARY KAFB, NM

.A

NATIONAL ADWSO~ COMMITTEEIllllmlllllllllllullllu

FOR AERONAUTICS ooLL27Cl.

TECHNICAL NOTE 32~

AERODYNAMIC CHARACTERISTICS OF THE

NACA 64-010 AND0010-1.10 @/l. 0~ AIRFOIL SECTIONS

AT MACH NUMBERS FRCM 0.30 TO 0.85 ANDREYNOLW

mms FW 4.o x 106 TO 8.o x 106

By Iaurence K. Loftin, Jr.

A short two-dimensional investigation has been made in the Langleylow-turbulencepressure tunnel to determine the aerodynamic character-istics of the NACA 64-o1o and 0010-1.10 40/1.051 airfoil sections. Theinvestigation covered a Mach nuniberrange from 0.30 to 0.85 and the corre-

sponding Reynolds number range extended from 4.0 x 106 to 8.o x 106. The

PWose of the invesii!+tionwas to determine the extent to which therelative mrits of the two airfoil sections, as indicated by previousinvestigations (NACA BM A9G18 and RM A9E31) at Reynolds nurtibersfrom

1.0 x ld to 2.0 x l&, might be alteredby ficreases in the Reynoldsnumber. The results indicated that the increment between the higher dragof the NACA 0010-1.10 40/1.051 airfoil section and the drag of the NACA64-o1o airfoil section shown by the data of NACA RM A9W8 and RM A9E31for moderate lift coefficients and relatively high subsonic speeds wasmuch sma12er in the present higher Reynolds nunber investi~tion.

INTRODUCTION

A recent airfoil-selectionproblem involved a choice between theNACA @+-010 section and a 10-percent-thick section of the mcdifiedNACA 4-digit-series family. A comparison of the data of reference 1for the modified 4-digit-series section with those of reference 2 forthe NACA @+-010 section indicated that the mmiified section had thehighest drag-divergence Mach number for the low lift coefficient, high-speed condition; whereas the drag of the NACA @l-010 section was lowerthan that of the modified section for the cruise condition at moderatelift coefficients= The investigations of references 1 and 2 were made.at relatively low Reynolds nunibers(R = 1.OX M$ -b 2.ox1o6), hOw-ever, which raised a question as to the effect of increasing the Reynolds

r<

.—— .— — — ~“

2 NACA TN 32ti

nuniberon the comparative drag characteristics of the two airfoil sec-tions. Since the relative merits of the two families of airfoils havebeen of considerable interest in the past fewyearsj a short two-dimensional tivestigationwas undertaken in the Langley low-turbulencepressure tunnel ti.determine the effect of Reynolds number on the dragcharacteristics of the NACA 64-o1o airfoil section and the 10-percent-thick, modified NACA 4-digit-series airfoil section. Lift, Wg, andpitching-moment data are presented herein for the two airfoil sections.The Mach nuuiberrange of the investigationwas from 0.30 to 0.85 and the

corresponding Reynolds nuniber=ge was from 4.0

x distance

Y distance

a section

Czsection

cd section

section

(

SYMBaLs

along chord

normal to chord

mgle of attack

lift coefficient

irag coefficient

pitching-momentcoefficient

R Reynolds number based on wing chordand densi@

M free-stream Mach numiber

APPARATUS AND mm

The two airfoils investigated were the NACA

X 1($ b 8.OX 106.

about quarter chord

and free-stream velocity

0010-1.10 40/1.051 andthe 64-o1o sections. The desib=tion NACA 0010-1.10 40/1.051 describesan airfoil which was obtained by modifying the NACA 0010 basic thicknessform. The 0010 appearing in the designation indicates a symmetricalairfoil of 10-percent thickness; the 1.10 indicates the size of theleading-edge radius in percent of the chord; the h indicates theposition of maximum thickness in percent of the chord; and the 1.051 isan index nuniberwhich is indicative of the trailing-edge angle with the

.,

___ ..—.— .—

NACA TN 3244 3

.

>

actual angle being twice the arc tangent of the prtiuct of the thiclmess-chord ratio and the index nuniberl.0~1. (It is perhaps of in~rest thatthe NACA ~10-1.10 40/1.051 section is similar to the well-known NACA0010-64 section except for a reduced trailing-edge angle.) Both modelswere of l-foot chord and were constructed of aluminum alloy. Ordinatesof the two airfoil sections are given in table I and sketches of thetwo profiles are shown in figure 1.

me tests were made in the Langley low-turbulence pressure tunnel

(ref. 3). The test section of the tunnel measures 3 feet by 7~feet

and was so designed that the models, when mounted in the tunnel, com-pletely spanned the 3-foot dimension. The model passed through slots inthe tunnel walls and was attached to the strain-gagebalance which wasemployed for lift, drag, and pitching-moment measurements. Iabyrinth-type seals were provided at each end of the model to minimize the effectof air leakage through the slots in the tunnel wall. Amore completedescription of the technique of making two-dhnensionalmeasurementswith the strain-gagebalance nmy be found in reference 4.

The tests of each model consisted in measurements of the lift, drag,and pitching moment for angles of attack from -2° to P and for a Mach num-ber range extending from about 0.30 to 0.85. An atmosphere of lZreon-12ata stagnation pressure of approxhately 24 inches of mercury absolute wasmaintained in the tunnel for all the tests. The relationship betweenReynolds number and Mach number for this stagnation pressure and a l?reonpurity of 95 percent by weight is shown in figure 2. lhe model surfaceswere polished to a high degree of smoothness at the time of model installa-tion in the tunnel. The drag coefficientsmeasured, however, probablydo not correspond to extensive regions of lsminar flow, since the use ofl?reon-12as a test medium makes unfeasible the almost continuous atten-tion to model surface condition which is required in order to maintainextensive lsminsr layers.

CORRECTIONS AND PRECISION OF MEASUREMENT

Corrections for wind-tunnel-wall effects have been applied to allthe data according to the methods of reference 5. Conversion of the dataobtained in Freon-12 to equivalent air data has been carried out by themethod of reference 6. The choking Mach nunibersin air were determinedfrom the measured values in llreonby the methods of reference 6 andvaried from about 0.875 to 0.850 as the angle of attack varied from 00to p.

4 NACA TN 3244

T& accuracy of the measured quantities has been estimated on thebasis of the limitations of the balance and is indicated for varioustest conditions in the following table:

M c1 cd %c/4

0.85 *0.001 *o.CQ03 *0.0002.60 2.002 *.0006 2.0003

● 35 **W * .ool~ t .0008L

RESUL’ISAND DISCUSSION

The basic data obtained in the investigation are presented in fig-ures 3 to 8 in the form of section lift, drag, and quarter-chordpitching-moment coefficient against Mach ntier for various angles of attack.Uft-drag polars obtained from cross plots of these data are presentedin figure 9 for several Mach numibers. A study of the drag polars givenin figure 9 indicates the same trends as shownby a comparison of thedata of references 1 and 2. The drag coefficients of the NACA h

0010-1.10 40/1.051 airfoil section appear in most instances to be some-what higher than those of the NACA 64-o1o airfoil section for Mach num-bers up to 0.75 and lift coefficients above 0.10. For Mach nunibersabove0.75, the advantage seems tobe with the NACA 0010-1.10 40/1.051 airfoilsection, at least for certain portions of the lift-coefficient range. Acomparison of the section drag characteristics of the two airfoil sec-tions as determined in the present tivestigation and in the investiga-tions of references 1 and 2 is protided in figure 10 in which sectiondrag coefficient is plotted against Mach nuniberfor section lift coeffi-cients of 0.2, 0.4, and 0.6. The most significant differences betweenthe results obtained in the investigations of references 1 and 2 andthose of the present higher Reynolds nuuiberinvestigation are seen in thedata for section lift coefficients of 0.4 and 0.6 at Mach numibersbelowthe force break. For these conditions, the present data indicate theincrement in drag betweenwith the increments shown

the two airfoils ~ be smaller in comparisonby the data of references 1 and 2.

CONCLUDING REMARKS

A short two-dimensional investigation has been made in the Langleylow-turbulence pressure tunnel to determine the aerodynamic character-istics of the NACA 64-o1o and 0010-1.10 40/1.O’jlairfoil sections. Theresults indicated that the increment between the higher drag of the NACACM)1O-1.1O40/1.051 airfoil section and the drag of the NACA 64-o1o air-foil section shown by the dab of NACA RM A5G18 and RM A9E31 for moderate *

,

.

NACA TN 32&

lift coefficients and relatively highin the present higher Reynolds number

Lan@y Aeronautical.Laboratory,

mibsonic speedsinvestigation.

5

was much smaller

National Advisozy Committee for Aeronautics,Iangley Field, Vs., February 24, 1954.

REFERENc@

1. Summers, James L., and Grahsm, Donald J.: Effects of SystematicChanges of Trailing-Edge Angle and Leading-Edge Radius on theVariation With Mach lhmiberof the Aerodynamic Characteristics ofa 10-Percent-Chord-!ChickNACA Airfoil Section. NACARMA9G18,1949.

2. Hemenover, Albert D.: Tests of the NACA64-O1O and @tAOIO AirfoilSections at High Subsonic lkch Numbers. mcARMA91Z51, 1949.

3. Von Doenhoff, Albert E., and Abbott, I’YankT., Jr.: The LangleyTwo-Dimensional Low-TurbulencePressure Tunnel. NACA TN 12@,1947.

4. Laftin, Laurence K., Jr., and Von Doenhoff, Albert E.: ExploratoryWvestigation at El@ and low Subsonic Mach Nunibersof TwoExperimental 6-Percent-ThickAirfoil Sections Designed To HaveHigh MsXbnum Iift Coefficients. NACAR.ML511’06, 1951.

5. Allen, H. Jtian, and Vincenti, Walter G.: Wall ~terference in aTwo-Dimensional-FlowWind Tunnel, With Consideration of theEffect of Compressibili@. NACA Rep. 782, 1~. (SupersedesNACAWR A-65.)

6. Von Doenhoff, Albert E., Braslow, Albert L., and Schwsrtzberg,Milton A.: Studies of the Use of Freon-12 As a Wind-TunnelTesting Medium. NACATN 3000, 1953.

—.._—- _.—. —— ——— - --—-—

NACA TN 3244

TABLE I

ORDINATES FOR I?ACA64-010 AND CO1O-1.10 40/1.051 AIRFOIZ SECTIONS

NACA @&OIO

X2percent chord

o

.5● 75

1.&2.557.5

101520

%

45

g

65707580859095

100

Y9percent chord

o

.820

.9@1.2501.7012.3432.8263.2=3.&24.3024.6394.8644.9804.9884.8434.5864.2383.8203.3452.8272.281_1.7221.176

.671

.2480

hE. radius: 0.720 percent chord

NACAOO1O-1.1O 40/1.051

x, Y>percent chord percent chord

o 0I.25 1.4662.5 I.9665.0 2.5897.5 3.00910 3.33715 3.845

4.240; 4.79140 5.00050 4.78360 4.19770 3.33880 2.30590 1.19395 .638100 .lci)

,.E.radius: 1.10 percent chord

— —

7

.

NACA 64-010

.

NACA 0010-1.10 40/1.051

Figure 1.. l%ofile shapes of NAC!A@&OIO and 0010-1.10 40/1.051 airfoil

sections.

.

,

8 NACA TN 32& .

.

9 x 106

08

)’

H 1

EJAIIII> II 111111u b

5

4

.

.3 .4 .5 .6 .7 .8 .9Mach number,M

Figure 2.- Approximte variation of Reynolds number with Mach number forstagnation pressure of 24 inches of mercury absolute and I?reonpuri@of 95 percent by weight.

,

A

NACA TN 324.4“9

.

l.lr 1 1

I- I--t”. .4LLLJ__L I

I 1 11 I v I l\ A

ii “n—H—H—M

0 3 . -cl- — -e- -o- ~ - 4

-. 1

-.2(> ?

-c- ~I

‘3-

-. /

I-. 4

.2 93 .4 .5 .6 ●7 .gMach number,M

.9

Figure 3.- Section lift coefficient of NACA 64-010 airfoil section as afunction of Mach number for variou.sangles of attack.

—.- —— — —--—.. —. . —

—-. ——. — —

10 NACA TN 3244

: ao:

02

DO

Vo

AO

o❑o

Figure 4.- Sectionfiction t

.3.2

.ll

.10

.09

.W

.07

.06

II III

I 1. I I A I I I III

I / II

.05

.04

.03

.Ce

.O1

?,~.9

Hachnumber,H

‘f

drag coefficient of NACA 64-010 airfoil section as a>f lkch number for v-ariousangles of attack.

.—

NACA TN 3244 U

.

$Eo

$

o

❑ 0

00

AO

Vo

DO

40

70

Figure 7.- Section quarter-chordpitching-nmment coefficient of NACA64-010 airfoil section as a function of I@ch number for various&es of attack.

— .—

12

Figure 6.- Setsection as

1.1.I 1a, deg

1.C. -0 -2

g: /’ ‘\

— -v 11?

.9 — - :? I A \ H

r r

.g— .

b‘/ ~

/t \

/b “

.7 4 ~ fl \< /

l\/

/ }

.6 / x KI

Y ‘u- ?

/

●5/ ‘.

/ // \y

v-/ ~

A.4 b \\

/

e * ~

.3 - $’- W

</

.2b

.1-

0 •1 i-l -5 — -0- + - -& ~

-. 1-

-. 2 — — + .—

n

I ~

-.3- 7 r/

-. 4-.2 .3 ●4 .5 .6 .7 .g .9

Mach namber,H

:tion lift coefficient of NACA 0010-1.10 hO/1.051 airfoila function of Mach numiberfor various angles of attack.

—.— ——— ...—— . ..— — . .

NACA TN 32~ 13

Do

Vo

AO

000

.14

.13

.12

.J_J_J_ v 7 ! I I I/ I II /l

.10

.09

.Og

.07

.06

.03

.04

.03

.02

.01

0.2 .3 .4 .5 .6 .7 .g .9

Maohnumber, H

Figure 7.- Section drag coefficient of NACA 0010-1.10 kO/1.section as a function of Mach number for various angles

051 airfoilof attack.

— — ——- .——-———————- ——— ——-—-——

14 NACA TN 32kk

I

a, deg

Mach number, M

Figure 8.- Section quarter-chord pitching-moment coefficient of NACA0010-1.10 40/lo051 airfoil section as a function of Mach numberfor various angles of attack.

,1

,

I

I

.> 0

>

lACA &U10

‘------- IIMAIMU2-1.1OkO/l.C51 i?>

‘ootion‘t Ooeffl~OntJ%

.*

.55

.CQ

.O1

0

.’%

.9

.a

.Q’3

.0?

.G-l

0

4-, -. 2 0 ,2 .& .6 .6 1.0 1.P

Figure 9.- Variation of section drag coefficient wLth section lift coef-

ficient for NACA &OIO and 0210-1.10 40/1.051 airfoil sections for

several Mach numbers.

G

16.

NACA TN 3244

●

Md

-5c0.

u0

f%

NACA6+010__----_ --–-NA~ oo10-1.1040/1.051

.08

.07

.06

.05

.04

.03

.02

.01

0

.11

.10

.09

.08

.0’7

.06

.05

.04

.03

.02

.01-.4 -.2 0 .2 .4 .6 .8 1.0 1.2

, .+ , , , , , ,+’I I i

. Section lift coefficient,Cz

Md

-2

. lC

.o~

.OE

.07

.06

.05

.04

.o~

.02

~ol

o

.11

.lC

.0s

.OE

.07

.Of

.05

.04

.02-. -. 2 0 .2 .4 .6 .8 1.0 1.2

Seotion liftcoefficient,~

.

,.

Fimxre 9.- Concluded.

— —..

ANACA m 32W 17

NACA @t-010

)

Langley

.—-— —— -_ _ NACA 0010-1.1040/1.091 R = 4.0

—— NACA (%-0101

Ames 1

lw-turbulence

x 106 to 8.o

pressuretunnel

x 106

x 3-1/2 ft. tunnel

— -- NACA 0010-1.10 40/1.051J R = 1.0 X 106 to 2.OX 106 (ref. 1 )

.04

.04 ‘

I

.03 (I

Cz = 0.6//

!/ /

/.’02 /

/ I

/+ . /

/ / // /.01 * = = — — — ‘

;;

ocl = 0.4

~-

— _

1’

Cz=() .2 ///1’t ,,

/—— —— —_. -— /

0 .02

.01

0. 3 .4 .5 .6 .7 .8 .9

Mach number, M

Figure 10.- Effect of Reynolds nuuiberon variation of section dragficient with Mach nuriberfor NACA @l-O1O-and 0010-1.10 kO/1.051foil sections..-

,

NAcA-L8ngley-8-.27-64-1000

coef-air-

..— ..—— .— —-_ . . .