Loftin, Laurence K , Jr

English

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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 Laboratory
Langley Field, Va
Washington
August 1954
.—. .,
I
—.--+ ___ .- ..<.. ..-~. . . . . . . . ,. .-.. . . .. . . . . . .. ---- ._. . . . . . . . . —>.- .—.
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TECH LIBRARY KAFB, NM
.A
NATIONAL ADWSO~ COMMITTEE
Illllmlllllllllllullllu
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 Langley
low-turbulencepressure tunnel to determine the aerodynamic character-
istics of the NACA 64-o1o and 0010-1.10 40/1.051 airfoil sections. The
investigation 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 the
relative mrits of the two airfoil sections, as indicated by previous
investigations (NACA BM A9G18 and RM A9E31) at Reynolds nurtibersfrom
1.0 x ld to 2.0 x l&, might be alteredby ficreases in the Reynolds
number. The results indicated that the increment between the higher drag
of the NACA 0010-1.10 40/1.051 airfoil section and the drag of the NACA
64-o1o airfoil section shown by the data of NACA RM A9W8 and RM A9E31
for moderate lift coefficients and relatively high subsonic speeds was
much sma12er in the present higher Reynolds nunber investi~tion.
INTRODUCTION
A recent airfoil-selectionproblem involved a choice between the
NACA @+-010 section and a 10-percent-thick section of the mcdified
NACA 4-digit-series family. A comparison of the data of reference 1
for the modified 4-digit-series section with those of reference 2 for
the NACA @+-010 section indicated that the mmiified section had the
highest drag-divergence Mach number for the low lift coefficient, high-
speed condition; whereas the drag of the NACA @l-010 section was lower
than that of the modified section for the cruise condition at moderate
lift 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 have
been of considerable interest in the past fewyearsj a short two-
dimensional tivestigationwas undertaken in the Langley low-turbulence
pressure tunnel ti.determine the effect of Reynolds number on the drag
characteristics of the NACA 64-o1o airfoil section and the 10-percent-
thick, modified NACA 4-digit-series airfoil section. Lift, Wg, and
pitching-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
Cz section
cd section
section
(
SYMBaLs
along chord
normal to chord
mgle of attack
lift coefficient
irag coefficient
pitching-momentcoefficient
R Reynolds number based on wing chord
and 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 and
the 64-o1o sections. The desib=tion NACA 0010-1.10 40/1.051 describes
an airfoil which was obtained by modifying the NACA 0010 basic thickness
form. The 0010 appearing in the designation indicates a symmetrical
airfoil of 10-percent thickness; the 1.10 indicates the size of the
leading-edge radius in percent of the chord; the h indicates the
position of maximum thickness in percent of the chord; and the 1.051 is
an 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 that
the NACA ~10-1.10 40/1.051 section is similar to the well-known NACA
0010-64 section except for a reduced trailing-edge angle.) Both models
were of l-foot chord and were constructed of aluminum alloy. Ordinates
of the two airfoil sections are given in table I and sketches of the
two 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 in
the tunnel walls and was attached to the strain-gagebalance which was
employed for lift, drag, and pitching-moment measurements. Iabyrinth-
type seals were provided at each end of the model to minimize the effect
of air leakage through the slots in the tunnel wall. Amore complete
description of the technique of making two-dhnensionalmeasurements
with 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-12at
a stagnation pressure of approxhately 24 inches of mercury absolute was
maintained in the tunnel for all the tests. The relationship between
Reynolds number and Mach number for this stagnation pressure and a l?reon
purity of 95 percent by weight is shown in figure 2. lhe model surfaces
were polished to a high degree of smoothness at the time of model installa-
tion in the tunnel. The drag coefficientsmeasured, however, probably
do not correspond to extensive regions of lsminar flow, since the use of
l?reon-12as a test medium makes unfeasible the almost continuous atten-
tion to model surface condition which is required in order to maintain
extensive lsminsr layers.
CORRECTIONS AND PRECISION OF MEASUREMENT
Corrections for wind-tunnel-wall effects have been applied to all
the data according to the methods of reference 5. Conversion of the data
obtained in Freon-12 to equivalent air data has been carried out by the
method of reference 6. The choking Mach nunibersin air were determined
from the measured values in llreonby the methods of reference 6 and
varied from about 0.875 to 0.850 as the angle of attack varied from 00
to p.
4 NACA TN 3244
T& accuracy of the measured quantities has been estimated on the
basis of the limitations of the balance and is indicated for various
test conditions in the following table:
M c1 cd % c/4
0.85 *0.001 *o.CQ03 *0.0002
.60 2.002 *.0006 2.0003
l 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 presented
in figure 9 for several Mach numibers. A study of the drag polars given
in figure 9 indicates the same trends as shownby a comparison of the
data 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 nunibersabove
0.75, the advantage seems tobe with the NACA 0010-1.10 40/1.051 airfoil
section, at least for certain portions of the lift-coefficient range. A
comparison 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 section
drag coefficient is plotted against Mach nuniberfor section lift coeffi-
cients of 0.2, 0.4, and 0.6. The most significant differences between
the results obtained in the investigations of references 1 and 2 and
those of the present higher Reynolds nuuiberinvestigation are seen in the
data for section lift coefficients of 0.4 and 0.6 at Mach numibersbelow
the force break. For these conditions, the present data indicate the
increment in drag between
with the increments shown
the two airfoils ~ be smaller in comparison
by the data of references 1 and 2.
CONCLUDING REMARKS
A short two-dimensional investigation has been made in the Langley
low-turbulence pressure tunnel to determine the aerodynamic character-
istics of the NACA 64-o1o and 0010-1.10 40/1.O’jlairfoil sections. The
results indicated that the increment between the higher drag of the NACA
CM)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 high
in the present higher Reynolds number
Lan@y Aeronautical.Laboratory,
mibsonic speeds
investigation.
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 Systematic
Changes of Trailing-Edge Angle and Leading-Edge Radius on the
Variation With Mach lhmiberof the Aerodynamic Characteristics of
a 10-Percent-Chord-!ChickNACA Airfoil Section. NACARMA9G18,
1949.
2. Hemenover, Albert D.: Tests of the NACA64-O1O and @tAOIO Airfoil
Sections at High Subsonic lkch Numbers. mcARMA91Z51, 1949.
3. Von Doenhoff, Albert E., and Abbott, I’YankT., Jr.: The Langley
Two-Dimensional Low-TurbulencePressure Tunnel. NACA TN 12@,
1947.
4. Laftin, Laurence K., Jr., and Von Doenhoff, Albert E.: Exploratory
Wvestigation at El@ and low Subsonic Mach Nunibersof Two
Experimental 6-Percent-ThickAirfoil Sections Designed To Have
High MsXbnum Iift Coefficients. NACAR.ML511’06, 1951.
5. Allen, H. Jtian, and Vincenti, Walter G.: Wall ~terference in a
Two-Dimensional-FlowWind Tunnel, With Consideration of the
Effect of Compressibili@. NACA Rep. 782, 1~. (Supersedes
NACAWR A-65.)
6. Von Doenhoff, Albert E., Braslow, Albert L., and Schwsrtzberg,
Milton A.: Studies of the Use of Freon-12 As a Wind-Tunnel
Testing 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
X2
percent chord
o
.5
l 75
1.&
2.5
5
7.5
10
15
20
%
45
g
65
70
75
80
85
90
95
100
Y9
percent chord
o
.820
.9@
1.250
1.701
2.343
2.826
3.2=
3.&2
4.302
4.639
4.864
4.980
4.988
4.843
4.586
4.238
3.820
3.345
2.827
2.281_
1.722
1.176
.671
.248
0
hE. radius: 0.720 percent chord
NACAOO1O-1.1O 40/1.051
x, Y>
percent chord percent chord
o 0
I.25 1.466
2.5 I.966
5.0 2.589
7.5 3.009
10 3.337
15 3.845
4.240
; 4.791
40 5.000
50 4.783
60 4.197
70 3.338
80 2.305
90 1.193
95 .638
100 .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
0
8 )’
H 1
EJAIIII> II 111111
u b
5
4
.
.3 .4 .5 .6 .7 .8 .9
Mach number,M
Figure 2.- Approximte variation of Reynolds number with Mach number for
stagnation pressure of 24 inches of mercury absolute and I?reonpuri@
of 95 percent by weight.
,
ANACA TN 324.4
“9
.
l.lr 1 1
I- I--t”
. .4LLLJ__L I
I 1 1
1 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 l7 .g
Mach number,M .9
Figure 3.- Section lift coefficient of NACA 64-010 airfoil section as a
function of Mach number for variou.sangles of attack.
—.-
—— — —--—.. —. .
—
—-. ——. — —
10 NACA TN 3244
: ao
:
02
DO
Vo
AO
oqo
Figure 4.- Section
fiction t
.3.2
.ll
.10
.09
.W
.07
.06
II II
I
I 1. I I A I I I I
II
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
.
$
E
o
$
o
q 0
00
AO
Vo
DO
40
70
Figure 7.- Section quarter-chordpitching-nmment coefficient of NACA
64-010 airfoil section as a function of I@ch number for various
&es of attack.
— .—
12
Figure 6.- Set
section as
1.1.
I 1
a, deg
1.C. - 0 -2
g: /’ ‘\
— -v 1
1?
.9 — - :? I A \ H
r r
.g
— .
b‘
/ ~
/
t \
/
b “
.7 4 ~ fl \
< /
l\/ / }
.6 / x KI
Y ‘
u- ?
/
l5
/ ‘.
/ /
/ \y
v-
/ ~
A
.4 b \\/
e
* ~
.3 - $’- W
<
/
.2 b
.1-
0 •1 i-l -5 — -0- + - -
& ~
-. 1-
-. 2 — — + .
—
n
I ~
-.3- 7 r
/
-. 4-
.2
.3 l4 .5 .6 .7 .g .9
Mach namber,H
:tion lift coefficient of NACA 0010-1.10 hO/1.051 airfoil
a 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 airfoil
of attack.
— — ——-
.——-——————- ——— ——-—-—
—
14 NACA TN 32kk
I
a, deg
Mach number, M
Figure 8.- Section quarter-chord pitching-moment coefficient of NACA
0010-1.10 40/lo051 airfoil section as a function of Mach number
for various angles of attack.
,1
,
II
.> 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
l
M
d
-5
c
0
.
u
0
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
M
d
-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.
—
—..
A
NACA m 32W 17
NACA @t-010
)
Langley
.—-— —— -_ _ NACA 0010-1.1040/1.091 R = 4.0
—— NACA (%-010
1
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 * = = — — — ‘
;;
o
cl = 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 drag
ficient with Mach nuriberfor NACA @l-O1O-and 0010-1.10 kO/1.051
foil sections.
.-
,
NAcA-L8ngley-8-.27-64-1000
coef-
air-
..— ..—— .— —-_ . . .


Aerodynamic characteristics 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 10(exp. 6) to 8.0 x 10(exp. 6)

Aerodynamic characteristics 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 10(exp. 6) to 8.0 x 10(exp. 6)

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