REPORT No. 460
THE CHARACTERISTICS OF
78 RELATED AIRFOIL SECTIONS FROM TESTS
IN THE VARIABLE-DENSITY WIND TUNNEL
By EASTMAN N. JACOBS, KENNETH E. WARD
and ROBERT M. PINKERTON
Langley Memorial Aeronautical Laboratory
REPRINT OF REPORT No. 460, ORIGINALLY PUBLISHED NOVEMBER 1933
27077 0-36-1
1
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
NAVY BUILDING, WASHINGTON, D.C.
(An independent Government establishment, created by act of Congress approved March 3,1915, for the supervision and direction of theselentific
study of the problems of flight. Its membership was increased to 15 by act approved March 2, 1929 (Public, No. 908, 70th Congress). It eonsista
of members who are appointed by the President, all of whom serve se, such without compensation.)
JOSEPH S. AMEs, Ph.D., Chairman,
President, Johns Hopkins University, Baltimore, Md.
DAVID W. TAYLOR, D. Eng., Vice Chairman,
Washington, D.C.
CHARLES G. ABBOT, Sc.D.,
Secretary, Smithsonian Institution, Washington, D.C.
LYMAN J. BRIGGS, Ph.D.,
Director, Bureau of Standards, Washington, D.C.
ARTHUR B. COOK, Captain, United States Navy,
Assistant Chief, Bureau of Aeronautics, Navy Department, Washington, D.C.
WILLIAM F. DURAND, Ph.D.,
Professor Emeritus of Mechanical Engineering, Stanford University, California.
BENJAMIN D. FOULOIS, Major General, United States Army,
Chief of Air Corps, War Department, Washington, D.C.
HARRY F. GUGGENHEIM, M.A.,
Port Washington, Long Island, New York.
ERNEST J. KING, Rear Admiral, United States Navy,
Chief, Bureau of Aeronautics, Navy Department, Washington, D.C.
CHARLES A. LINDBERGH, LL.D.,
New York City.
WILLIAM P. MACCRACKEN, Jr., Ph.B.,
Washington, D.C.
CHARLES F. MARVIN, Sc.D.,
Chief, United States Weather Bureau, Washington, D.C.
HENRY C. PRATT, Brigadier General, United States Army.
Chief, Mat€riel Division, Air Corps, Wright Field, Dayton, Ohio.
EDWARD P. WARNER, M.S.,
Editor "Aviation," New York City.
ORVILLE WRIGHT, Sc.D.,
Dayton, Ohio.
GEORGE W. LEWIs, Director of Aeronautical Research.
JOHN F. VICTORY, Secretary.
HENRY J. E. REID, Engineer in Charge, Langley Memorial Aeronautical Laboratory, Langley Field, Va.
JOHN J. IDE, Technical Assistant in Europe, Paris, Franed.
EXECUTIVE COMMITTEE
JOSEPH S. AMEs, Chairman.
DAVID W. TAYLOR, Vice Chairman.
WILLIAM P. MACCRACKEN, Jr.
CHARLES F. MARVIN.
HENRY C. PRATT.
EDWARD P. WARNER.
ORVILLE WRIGHT.
CHARLES G. ABBOT.
LYMAN J. BRIGGS.
ARTHUR B. COOK.
BENJAMIN D. Fo ULOIS.
ERNEST J. KING.
CHARLES A. LINDBERGH.
JOHN F. VICTORY, Secretary.
E
REPORT No. 460
THE CHARACTERISTICS OF 78 RELATED AIRFOIL SECTIONS FROM TESTS IN THE
VARIABLE-DENSITY RIND TUNNEL
By EASTMAN N. JACOBS, KENNETH E. WARD, and ROBERT M. PINKERTON
REPRINT OF REPORT No. 460, ORIGINALLY PUBLISHED NOVEMBER 1939
SUMMARY
ence 1) but, with the exception of the M-series and a
An investigation of a large group of related airfoils series of propeller sections, the airfoils have not been
was made in the N.A.C.A. variable-density wind tunnel systematically derived in such a way that the results
at a large value of the Reynolds Number. The tests were could be satisfactorily correlated.
The design of an efficient airplane entails the careful
made to provide data that may be directly employed for a
rational choice of the most suitable airfoil section for a balancing of many conflicting requirements. This
given application. The variation of the aerodynamic statement is particularly true of the choice of the wing.
characteristics with variations in thickness and mean-line Without a knowledge of the variations of the aerodynamic characteristics of the airfoil sections with the
form were therefore systematically studied.
The related airfoil profiles for this investigation were variations of shape that affect the weight of the strucdeveloped by combining certain profile thickness forms, ture, the designer cannot reach a satisfactory balance
obtained by varying the maximum thickness of a basic between the many conflicting requirements.
The purpose of the investigation reported herein was
distribution, with certain mean lines, obtained by varying
the length and the position of the maximum mean-line to obtain the characteristics at a large value of the
ordinate. A number of values of these shape variables Reynolds Number of a wide variety of related airfoils.
were used to derive a family of airfoils. For the purposes The benefits of such a systematic investigation are
of this investigation the construction and tests were limited evident. The results will greatly facilitate the choice
to 68 airfoils of this family. In addition to these, several of the most satisfactory airfoil for a given application
supplementary airfoils have been, included in order to and should eliminate much routine airfoil testing.
study the effects of certain other changes in the form of the Finally, because the results may be correlated to
indicate the trends of the aerodynamic characteristics
mean line and in the thickness distribution.
The results are presented in the standard ,graphic form with changes of shape, they may point the way to the
representing the airfoil characteristics for infinite aspect design of new shapes having better characteristics.
Airfoil profiles may be considered as made up of cerratio and for aspect ratio 6. A table is also given by
means of which the important characteristics of all the tain profile-thickness forms disposed about certain
airfoils may be conveniently compared. The variation of mean lines. The major shape variables then become
the aerodynamic characteristics with changes in shape is two, the thickness form and the mean-line form. The
shown by additional curves and tables. A comparison thickness form is of particular importance from a
is made, where possible, with thin-airfoil theory, a structural standpoint. On the other hand, the form of
the mean line determines almost independently some
summary of which is presented in an appendix.
INTRODUCTION
The forms of the airfoil sections that are in common
use today are, directly or indirectly, the result of
investigations made at Gottingen of a large number of
airfoils. Previously, airfoils such as the R.A.F. 15
and the U.S.A. 27, developed from airfoil profiles
investigated in England, were widely used. All these
investigations, however, were made at low values of
the Reynolds Number; therefore, the airfoils developed
may not be the optimum ones for full-scale application.
More recently a number of airfoils have been tested in
the variable-density wind tunnel at values of the
Reynolds Number approaching those of flight (refer-
of the most important aerodynamic properties of the
airfoil section, e.g., the angle of zero lift and the
pitching-moment characteristics.
The related airfoil profiles for this investigation were
derived by changing systematically these shape variables. The symmetrical profiles were defined in terms
of a basic thickness variation, symmetrical airfoils of
varying thickness being obtained by the application
of factors to the basic ordinates. The cambered profiles were then developed by combining these thickness
forms with various mean lines. The mean lines were
obtained by varying the camber and by varying the
shape of the mean line to alter the position of the
maximum mean-line ordinate. The maximum ordinate
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
of the mean line 2s referred to throughout this report as the
If the chord is taken along the x axis from 0 to 1,
camber of the airfoil and the position of the maximum
ordinate of the mean line as the position of the camber.
An airfoil, produced as described above, is designated by
a number of four digits: the first indicates the camber in
percent of the chord; the second, the position of the camber
in tenths of the chord from the leading edge; and the last
the ordinates y are given by an equation of the form
f y = ad ^x + ax + a2 x2 + a3x' + a4x4
The equation was adjusted to give the desired shape
by imposing the following conditions to determine the
constants:
two, the maximum thickness in percent of the chord.
(1) Maximum ordinate 0.1 at 0.3 chord
Thus the N.A.C.A. 2315 airfoil has a maximum camber
of 2 percent of the chord at a position 0.3 of the chord
from the leading edge, and a maximum thickness of 15
percent of the chord; the N.A.C.A. 0012 airfoil is a
symmetrical airfoil having a maximum thickness of 12
percent of the chord.
In addition to the systematic series of airfoils,
several supplementary airfoils have been included in
order to study the effects of a few changes in the form
of the mean line and in the thickness distribution.
x=0.3
y=0.1
dy/dx = 0
(2) Ordinate at trailing edge
x=1
y=0.002
(3) Trailing-edge angle
x=1
dy/dx= —0.234
bered airfoils, the 43 and 63 series (reference 4), the 45
(4) Nose shape
X=0.1
y = 0.078
The following equation satisfying approximately the
above-mentioned conditions represents a profile having
a thickness of approximately 20 percent of the chord.
and 65 series (reference 5), the 44 and 64 series (reference 6), and the 24 series (reference 7).
t y = 0.29690. 1/ — 0.12600x — 0.35160x 2 + 028430x3
— 0.10150x4
Preliminary results which have been published in-
clude those for 12 symmetrical N.A.C.A. airfoils, the
00 series (reference 2) and other sections having different nose shapes (reference 3); and those for 42 cam-
./
- ^^olo^°
a
-
N. A. C. A. family
t_e
r
O
^e
+ Clark Y
o Gdtt. 398
/0
./;
.2
.4
.5
.6
.7
.8
.9
/.O
0.12600 x -0.35160x' +0.28430 x' -0.10150x'
Basic ordinates of N.A.C.A. family airfoils (percent of chord)
±v = 0.29690 Arx_-
Ord____.I
5.0
1
O 1 31 157 4.3581 5.9261 7.0001 10.8051 &9091 9.6031 9.002110.0031 49.0721 8.823107 .sod 1 7d u, 1 84.8721 92.4131 01.344 1 1 0. 210
L.E. radius, 4.40.
FIGURE
I. Thickness variation
The tests were made in the variable-density wind
tunnel of the National Advisory Committee for Aeronautics during the period from April 1931 to February
1932.
DESCRIPTION OF AIRFOILS
Well-known airfoils of a certain class including the
Gottingen 398 and the Clark Y, which have proved to
be efficient, are nearly alike when -their camber is
removed (mean line straightened) and they are reduced
to the same maximum thickness. A thickness variation
similar to that of these airfoils was therefore chosen for
the development of the N.A.C.A. airfoils. An equation
defining the shape was used as a method of producing
fair profiles.
This equation was taken to define the basic section.
The basic profile and a table of ordinates are given in
figure 1. Points obtained by removing the camber
from the Gottingen 398 and the Clark Y sections, and
applying a factor to the ordinates of the resulting
thickness curves to bring them to the same maximum
thickness, are plotted on the above figure for comparison. Sections having any desired maximum thickness were obtained by multiplying the basic ordinates
by the proper factor; that is
±y,=6-t (0.29690.1/x-0.12600x-0.35160x2
+0,28430e-0.10150x')
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL
where t is the maximum thickness. The leading-edge
radius is found to be
and
yr=(1 p)2 [(1-2P)+2Px-el
a
=' .Iota
r` 2 0 20 ae)
(aft of maximum ordinate)
When the mean lines of certain airfoils in common
use were reduced to the same maximum ordinate and
compared it was found that their shapes were quite
different. It was observed, however, that the range
of shapes could be well covered by assuming some
simple shape and varying the maximum ordinate and
its position along the chord. The mean line was,
therefore, arbitrarily defined by two parabolic equations of the form
yo= be+ b i x+ b2 X2
where the leading end of the mean line is at the origin
and the trailing end is on the x axis at x=1. The
The method of combining the thickness forms with
the mean-line forms is best described by means of the
diagram in figure 2. The line joining the extremities
of the mean line is chosen as the chord. Referring to
the diagram, the ordinate yt of the thickness form is
measured along the perpendicular to the mean line
from a point on the mean line at the station along the
chord corresponding to the value of x for which yt
was'computed. The resulting upper and lower surface
points are then designated: values of the constants for both equations were then
where the subscripts u and l refer to upper and lower
surfaces, respectively. In addition to these symbols,
the symbol 0 is employed to designate the angle between the tangent to the mean line and the x axis.
This angle is given by
expressed in terms of the above variables; namely,
(1) Mean-line extremities
X=0
y^=0
X=1
y'=0
(2) Maximum ordinate of mean line
6 = tan-1 dx
x=p (position of maximum ordinate)
y Oa
lxa. ya
./0-
-
Stations x„ and x,
Ordinates y„ and yt
B=Ion-'
9 yt
dx
v Mean /ins
P1
p
Chord
^
Or /xt , yt/
Rodius fhrough end of chord
-./O L
x
x„ = x - y, sine Yu ' Y + yt cos B
xt = x + y, sin a yt = yt - yt cos e
/.00
Sample calculations for derivation of N.A.C.A. 6821
z
0
0.01250
1
.3000
.am
v,
ue
0
0
0.03314
0.00489
.10503
.07988
.0221
0
.06000
.04898
tan a
10.40000
sin v
aos a
0.37140
0.92840
.38333
.35793
0
-.07347
0
-.07327
-.17143
-.16897
y, sin e
0
0
.93375
0.01186
0.03094
.99731
0
-.00585
.10503
.07906
1
.98662
-.0037
z,
y. cos a
v.
........................
.00218
x,
0
m
0
0.00084
0.03683
0.01438
-0.02805
.30000.
.80585
.16503
.12863
.30000
.59415
.0218
.99963
-.04503
-.0307
1.00037
-.0218
Slope of radius through and of chord.
FIGURE 2.-Method of calculating ordinates of N.A.C.A. cambered airfoils.
yo=m(maximum ordinate)
dye/dx = 0
The resulting equations defining the mean line then
became
Myr=
p2
I2Px - el
(forward of maximum ordinate)
The following formulas for calculating the ordinates
may now be derived from the diagram:
xa =x-y, sin 6
ya =y,+y, cos 0
x l =x+y t sin 0
y t =yr -y, cos 0
Sample calculations are given in figure 2. The center
for the leading-edge radius is placed on the tangent to
the mean line at the leading edge.
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
A family of related airfoils was derived in the manner
described. Seven values of the maximum thickness,
0.06, 0.09, 0.12, 0.15, 0.18, 0.21, and 0.25; four values
of the camber, 0.00, 0.02, 0.04, and 0.06; and six values
of the position of the camber, 0.2, 0.3, 0.4, 0.5, 0.6, and
0.7 were used to derive the related sections of this
family. The profiles of the airfoils derived are shown
collectively in figure 3.
For the purposes of this investigation the construc-
tion and tests were limited to 68 of the airfoils. Tables
of ordinates at the standard stations are given in the
figures presenting the aerodynamic characteristics.
These ordinates were obtained graphically from the
computed ordinates for all but the symmetrical sec06
2206
2209
-G
X09
001Z- '— 2212_
2215
O55
C
L
0-018
-^
models, which are made of duralumin, have a chord
of 5 inches and a span of 30 inches. They were constructed from the computed ordinates by the method
described in reference 8.
Routine measurements of lift, drag, and pitching
moment about a point on the chord one quarter of the
chord behind its forward end were made at a Reynolds
Number of approximately 3,000,000 (tank pressure,
approximately 20 atmospheres). Groups of airfoils
were first tested to study the variations with thickness,
each group containing airfoils of different thicknesses
but having the same mean line. Finally, all airfoils
having a thickness of 12 percent of the chord were
2406
2306
2409
2309
2312 ^2412
X315
2318
2418
_.^1 '^
-2421
^_ C_^
^'
2218
^----2,506
2606
2509
2609
2612
2706
2709
-^2712
2615 -^
8 ' ^
2618
2918
-^
^ 4206
4306
4406
4506
4606
4706
4209
4309
4409
4509
4609
4709
9212
4312
4412
4512
4612
4712
4215
4315
4415
4515
4615
4715
4218
4318
4418
4518
4618
4718
4221
4321
4421
4521
4621
9721
—
6206
6306
6406
6506
6606
6706
6209
6309
6409
6509
6609
6709
6212
6312
6412
6512
6612
6712
6218
6318
6418
6518
6618 ^-6718
6221
6321
6421
6521
6621
0025
F—E
6721
3.—N.A.C.A. sWoll pmffiw.
tions. Two sets of trailing-edge ordinates are given.
Those inclosed by parentheses, which are given to
facilitate construction, represent ordinates to which
the surfaces are faired. In the construction of the
models the trailing edges were rounded off.
Three groups of supplementary airfoils were also
constructed and tested. The derivation of these airfoils will be considered later with the discussion.
APPARATUS AND, METHODS
A description of the variable-density wind tunnel
and the method of testing is given in reference 8. The
tested to study the variations with changes in the
mean line.
RESULTS
The results are presented in the standard graphic
form (figs. 4 to 80) as coefficients corrected after the
method of reference 8 to give airfoil characteristics for
infinite aspect ratio and aspect ratio 6. Where more
than one test has been used for the analysis, the infinite
aspect ratio characteristics from the earlier test have
been indicated by additional points on the figure. Table
I gives the important characteristics of all the.,airfoils.
T
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL Lw0r.
/2
0
U i 0
yWy p /0
.947
- /.307
/.777
-241
O
-2.673
20
40
60
80 100
Per cent ofch-d
-2.669
2.
.00/
-2.902
a
7
-2.282
,832
- /.3/2
- .724
.4031
-0.40
u
0
1
•l0
.09
.44
20 .40 G.OB
,y
1.8 .36 u .07.
L6 .32 8.06
32
28 ^
.O
24 0
V
L4 .26% 0.05
.v
28 0
LD
24 0
y
36 v
l
c. p.
20 w
16,
12^
CIE
8 ^^
/.2V .24 ^.04
v
k
K
40 20 °.03
u
.8 8 .16 0 .02
O
o
.0/
60./2
0
.4^ .0B
N 20^
C
0l 16 y
l
o /2^
g0
0 4u
.2
.04
v
oQ
0
Airfoil: N.A.C.A. 0006 R.N.:.^21Q000
Ve%(0../sec.): 66.5 _' 2
Sze: 5"x30"
Pres.(sthd.. otm.):20.8 Dofe: l-4-32
Where felted.L.MA.L. Test.• U.D.T. 744 -.4
Corrected for tunnel-wo ll effect.
0
.0
-4 W
-9 0
4
0
1216
8
20
4,^
u
0
0
-4 0
m
^-.z
.3
-.4
..4
24 28 32
Angle of .,tack, a (degrees)
-8 a
: i
R.N-
Airfoil: N.A.C.A. 0006
Test: V L
Date.-I-4-32
Corrected to infinite aspect ro
-/2
-/6
2 .4 .6 8 10 1.2 1.4
Lift coefficient, C
F-- 4.-N.A.C.A. 0000.W.H.
cu
sta. Op'r. C.'r.
U
v a /0
0
O O
125 1.420 -1. 961 U -C O
-1.96/ U
25 ,96/
5.0 666 -2666 4 0 -/0
7.5 3./50 -3.150
O 20 40 60 80
10
.5/2 -3.512
15 4.009 -4.009
Per cent ofchord
20 4.303 -4.303
25 4.456 -4.456
304.50/ -4.501
40 352 4.352
503. -' -3.97/
60 3.423 -3423
D
2.746
70 2.746
60
/.967 -/.967
901. 066-1.066
SU
.09
h
36 w
20 .40 ^'.. 08
32 ag
.44
l
N
1.8 .36 .g,.07
ea
1.6 .32 0.06
24
e U
1.4 .28- 0.05
20 w
100!095) (A 5J
/00 0 0
L.E. Rod.: 0.69
41
28
C
1
1?
c. p.
2421
q 2004/
1.2
t
/6 p
v
/2
8'^
.24 v.04
^
/.0 8.208 Q.03
u
.88.160 .02
0
80
.0/
. 6 .12
LD
o /6D 6i
^
a /2^Bi
8 0 /0,
° 4u
.o
04
G
l
k
4^
w
•4`1 .08,0
11
vi
C
0 0
c
-8 u
-6
-4
0
4
6
12 16
-4 0
.2 .04
Airfoil: N.A.CA. 0009 R.N.:3,2/0,000
Ve/.(Alsec.):68.5
Size: 5-x30"
Pres.(stnd.otm.):20.8 bbafe: /-6-32 -.2
Where tesled:L.M.A.L. Test:VOT 746 -.4
Corrected for tunnel-wall effect.
-4 F
Op
I
20 24 28 32
Angle of oftack, a (degrees)
"a ec
y -.2
°u
^ -.3
0 -.4
-4
Fla". 6.-N.A.C.A. 0000 MIMI.
Airfoil.• N.A. C. A. 0009
Dote. /-6-32
R. N.: 32/0,000
Test: V.D.T. 746
Corrected to inf)iVM aspect ratio
0 .2 .4 .6 .8 10 /.2 14
Lift coefficient. Cl
8
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAIITICS
20
Sla Up'r.
0
0
0
/.25 /.8.94-1.894
5.03.26/5
555 -2615
2.5
-3555
7.5 4.200 -4.200
v o /o
ut
o
/0 4.683 -4.683
/55.345 -5345
O
20 5738 -5.738
25 5.941 -5.941
306.002 -6.002
40 5.603 -5603
50 5.294 -5294
604.563-4.
70 3664 -3.664
563
80 2623 -2623
901.445- 1.446
95 .807 - .607
%00 (. ^) (-. p 6)
L. E. Rod.: 1.58
O
t
'w
a
v
./2
O
20 40 60 60
Per centofchord
00
2.0 .40 -08
N
/.8 .36 x.07
1.6 .32 0.06
C
C1 U
/.4 .281 0.05
,m V
26 0 0
24 0 20 e.p.
-0 40
L/
p /660
v.04
v ^=
1.01.208 .03
A
.16 02
o O
..6'.12
12 80
9 -/00
.4'.08
° 4uu
0Q
0
0 0
Airfoil: N.A. C.A. 0012 R.N:3,230,000
Size: SWO"
V.I.(k/sec.f: 68.4
Pres. (sfnd. otm.):207 Date. , 12-30-31
Where tested.• L.M.A.L. Test., V.D.T. 743
Corrected for funnel-wol/ effect.
c
-4 4
.0/
i
.2 .04
o
-8 u
./0.09
.44
u .2
u
-.3
+^
-.2
-.4
-8. -4 0 4 8 /2 16 20 24 26 32 Angle of oftoch, o: (degrees)
-.4
` -4
.4
.6
8
1.0
/.2
Lift coefficient. C.
FI°URE 0.-N.A.C.A. 0012 Oit 0.
Slo Up'r.
L'w'r
0 0
0
[252.367 -2367
2.53.268
5.0 4.443
7.5 5.250
/05853
/5 6.68/
207./72
25 7.42
7.
-3268
-4.443
-5.250
-5653
-6.681
-7./72
- 7.427
-7.502
-7.254
-6.618
-5.704
-4.580
a 20
/0
UNlp 0
0
./2
11
0
20 40
60
80
Per cent ofchord
00
30 50
40 7.254
50 6.618
60 5.704
70 4.S80
BO 3.279 -3279
90 /.8/0 -/.8/0
U
0
U
9
`o
/
O
$
.09
20 .40 r 08
C
v
/.B .36 `.07
/.006 -/.008
(-. 1561
l00 (./SB)
10
.44
V
1.6 .32 0.06
0
L. E. ROd.: 248
C
0 u
280 0
1.4 .26- 0.05
c. .
24 0 20
w
g20 40
L
.v
42
a /2 80
8,0100
0 4v
u
l
C
-4 0.
-8 u
Airfoih
.A. 0015 R.N.:3,200,000
Ve/.(ft./sec.): 68.4
fm):2/.O Dote:/2-9-31
.• LMAL. Test: VD.T. 728
For tunnel-wall effect
t
.2 .04 j-./
v
O
0 y -.2
u
-.2
3
-,4
0-.4
-6 -4 0 4 6 12 16 20 24 28 32 Angle of ot/oc/r, of (degrees)
.24•^
wy.04
,.0 208 .113
.8 w .16 0 .02
o a
.6 u .12
.0/
v
.4 v .08
.0
/
0 /6 60
l
^
.g
V
4
Lift coefficient C
Flo- 7.-N.A.C.A. 0015 etrlall.
1.4
9
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE -DENSITY WIND TUNNEL
20
c U /O
^c O
Co
Z53.9,11", 2 -.?.922
SO 5.332 -5.332
to 6.300 -6.300
/0 7,024 -7.024
20 6.0/8-_6.0/6
20 66 6.606
2569/ -6.9/2
30 9.003 -9.003
40 6. 041 -6.705
507.84/ -7.94/
60 6845-6.845
70 5436 -5.496
0
44.
40 60 BO /0
Per cent ofchord
0
0
24 L. 20
40
/.8 .36 x.07
32
28 ti
46 .32 0.06
,O
24 p
l
l
1 U
20
1.4 .28c 0.05
c. p.
y20x40
L/D
o /6A 60
l
01280
8 0100
0 4m
.o l
0Q
-,1'Cry
u
/6 p
/.21.24 x.04
M
/.Oy.20^4.03
a
U
.8.160 .02
.6 u .12^ .0/
q0
.4 -4 .08
a
-8
08
G
C
26
u
360)
2.0 .40
C
.2 .04
0
0
C,
Airfoil:N.A.CA.00/8 R.N.:,1150.000
Ue(.(H./sec.): 68.8
Size: S"x30"
Pres.(stnd. atm.): 20.9 Date: 1-6-32
/2 ig
8 l
k
4^
Op
-4 0
v
-8"'
-.2
3
Where tesfed:L.M.A.L. Test.• VD.T, 747
Corrected for funnel-wolf effect.
-.4 0-.4k
-8 -4 0 4 8 12 16 20 24 28 32 -4
AnO/e of oitock, a (degrees)
Fionaa 8 .-N.A.C.A. 0018 airfofl.
Airfoil: N.A.C,A. 00/8
Dote: / - 6 -32
Corrected to infinite a4
2
.4
.6
.8
/.0
cient.0
Lift coeffi
St. up'r. L'wr.
C C /O
l 0 33/4 -3.3/4 u s° 0
25 4.57 -4.576
k _/0
506.22/ -6.22/ y o
7.5 7.350 -7.350
0
l0 6./95 -6./°5
/59.354-9.354
20 to 1 -/0.04/
25 397 -/0397
30 5 -/0.503
40/0/5 -/0/55 50 9.26 -9.265
60 7.966 -7.986
80410
70 6.4/2 -6.9/2
-4590
90 2533 -2.533
95 /.4/2-/.4/2
lO
u0
O
/
N
.o
0
C
-8 u
80
t0
.44
R. .40
/.8 .36
/.6 .32
1
14 .28c
.v
1.21.24 1
v
/.0y.20°u
u
C
c.p.
tE
L/D
s
.l6 0
o O
.60
t: ./2
.4-'.08
.2 ,04
a
4u
l
C Q
b'0
O
L. n Had.: 4.65
0
24 a 20
k
V20 0 40
/6t 60
l
/2480
60(00
0
40
1011(.221,1-.2211
ac
2B 0
20.
Per centofcho rd
Airfoil: N.A.C.A. 0021 R.N, 3,190.000
t/ei.(H./sec.): 68.6
Size: 5-"x30"
Pres.(stnd.atm.):20.8 Date:1-7-32
Where tested'L.M.A.L. Test: VD.T. 748
Corrected for tunnel-wall effect.
0
0
-2
-.4
-8 -4 0 4 6 /2 16. 20 24 26 32
Angle of attack, of (degrees)
FM
27077 0-35--2
1
.09
44
L.E,Had: 3.56
v
48
/2
v ^ _/0
60 3935 -3935
90Z/ 72 -2.172
95 1.210 -/.2/
(./089) (-.0 9)
AX
0
0
a
z
>a Up'r. L'w'
10 264/ -2.841
9.-N.A.C.A. 0021 al/3olL
Lift coefficient,
io
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
Up'r. L'wi:
z D /0
d
V1R
2S 5^7 -5497
5.0 7.406 - 7.406
7S 6.750 -6.750
/0 9.756-9.756
/5//./36-/1./36
20//.953-/1.953
25/2.378 -/2.378
30 2504 -/2 504
40 2.09 -209
50/1.029-/1.029
60 9.507-9.507
70 7.633-7.633
80 5.465 -5465
90 30/6 -30/6
1.680- /.68
100 (.262) f-.262
0
0
0
LC. Rod.: 6.86
l
O
t
U
O
U
c
v
28
0
20 40 60 BO /L
Per centofchord
.44
2.0 .40
1.8 .36
1.6 .32
G°
/.4 .281
r
1
3
24 0 21
k
q2004(
16 61
0 /281
B o /Of
ti0 4 m
.o
C o-/0
O
IC.
c. p. 1
r-'
L 2,j. 24
11
v
1.0 ,y .20
A.16
v./6^
0 0
.6 u./2
.4 08
.2 .04
C
u
l
R 04
0
Airfoil: MA.CA. 0025 R.N.:3,230,000
Size: 5"x30 ••Ve/(fl/sec.): 68.3
Pres. (sfnd ofm.J. • 20.9 Date: 1-8-32
Where tested L.MA.L. Test: VD.T. 749
Corrected for tunnel-wall effect
.0
4 0.
V
-8
8
-4
0
4
8
12
16
20
24
An of attack, a (degrees)
28
0
-.2
-.4
32
Lift coefficient,
0026 airfoil.
1'..30.-N.A.O.A.
5/a.
0
a
g
O
a
W
28
0
Up'r. LWr.
O
252.44 -/.46
25 3.35 -/.96
5.0 4.62 -235
7.5 5.55-2
/O 6.27 -3/9
/57.25
-3.44
20 774 -3.74
25 793 -3.94
30 7.97 -4.03
40 7.66 -392
50 7.02 -3.56'
60 6.07 -3.05
70 4.90 -2.43
60 352 -1.74
90 /.93 - .97
95 1.05 - .ss
(./3) (-613)
00
L.E. Rad: 1.56
5tope
of
-d-0
V t 0
l
p -/0
R
O
20 40 60 60 11.
Per centofchord
44
2.0 .40
D
./0
40
.09
36
.08
32
.07
28
24
20
w ^
2j .24 v .04
16
v
101 20 ^ o .03
B ` .16 0
l
O
0 12 80
/2
.02
8'
O
6u./2
4
4".08
8'-/00
k c
0 4v
0
p
0 0 w _.2
-8
a
u.
l
o
48
44
U0 .06
1.4 .28 0x .05
c.p
L/
2
ll
d
/.8 .36
1.6 .32
0
24 ,3b 20
020040
l
p /6^ 60
l
D 20
y o !0
.2 .04
C
R 0Q
Aif.il.-N.A.C.A. 2212 R.N:3,220,000
Size: 5"k30"
V IMIsec.):68.4
P e5.(5ti7dolm.):208 Dote:12-2-31
Where tested:L.MA.L. Test.-VDT 719
j Corrected for tunnel-watt effect
_4 c0.
u
-8
-8
-4
0
4
6
/2
16
20
24
Angle of ottack, a (degrees)
28
0
-.2
`c
_4
32
0
.3
12
Airfoil.-NA.CA. 2212
R.N:3,220,000
Dofe:/2-2-3/
Test: VD.T. 7/9 -lB
Corrected to infinite ospecfrot/o
-. 4 -.4
-2
0
.2
.4
.6
.8
/.0
Lift coeff/cient,C
FIGME 11.-N.A.C.A. 2212 airfoll.
/.2
/.4
t.6
18
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE -DENSITY WIND TUNNEL
5^ up'r. L'^ r.
D /0
° g 0
ly
^ 0-/0
Z25 1.16 - .73
25 /.70 -.95
5.0 2.43 -/./5
7.5 30/ -/.22
2
20 40 60 80 /0
Per centofchord
0
/0 3.46 -1.22
/5 4.18 -/./6
20 4.65 -1.09
25 4,91 -1.04
30 S.00 -1.00
404.86-.94
50 4.49 - .8/
s0 3.92 - .65
70 3.19 - .48
60 230 - .33
90 /.26-.19
95 66 - .13
l00 (.06) (-.06)
0
U
0
c
/00
28
0 0
3
24 a 20
-
1
.44
48
/
44
/0
40
.09
360
.06
32 D
l
v
/.8 .36 .9 .07--28
C.E. Rod.:0.40
Slope afrod'us
lhraugh end of
chord: 2115
1.4 .28 P .OS
00,
.04
/6 0
z
C3
0
N
N O
m '^
/O.w.20$
C
LD
03
.02-a
eo./sa
.6^./2
v
u 60
U
0
0^
O
-4 0
v
Y
.2 .04 V`
Ca
0 0 v
0
c .3
1
.c
-4 0.
-8 u
-4
-6
-4
Airfoil: N.A.C.A. 2306 R.N.:3,080,000
Vel(ft./ser,4:69.8 -.2
Size; 5"X30"
Pres.(sfnd.otmJ206 Dofe:3-24-31
Where 1-1.1. L,M.A,I, Test: l<D.T. 660
Corrected for tunnel-wall effect -.4
0
4
6
12
16
20
24
An of attack, a (degrees)
28
F
Airfoi/. N.A.CA. 2306
R.N.:3080,'00 -/2
_ q9ote:9-24-31
Test:. V. D. T. 680 _/6
Corrected to infinite aspect ^-otio
-4 .2 0 .2 .4 .6 .8 10 /.2 1,4 /.6 18
32
Lift coefficient.0
FIGURE 12.-N.A.C.A. 23N airfoil.
^r.
St.
up'r, L'
I25 1.69 -1.16
2.5 2.39 -/.56
5.0 336 -2.0/
7.5 4. 9 -224
to
a. 7 -2.36
15 5.54 -250
m^ l0
U t 0
y _/0
C
0 20 40 60. 80 /0 7
Per centofchord
20 6.06 -2.52
25 6.37 -251
30 6.50 -250
40 632 -2.39
50 5.62 -2.13
60 5.07 -1.76
70 4./1 -1.36
80 2.96 - ,97
90 /.64 - .54
95 .66 - .33
100 (./01 /-./O)
loo - 0
L.E.ROd.:0.69
O
U
0O
ac
W
.44
R. C
oJrod'us
Through end of
280 0 5""'
chord: 2175
24 0 20
'20 Q 40
$1660
U
0^
.0
-q 0.
-8
/.8
.36
1.6
.32
V
/.4 .261
.v
1
L/D
1.0 y.20
u
8w
./6o
0 0
.6,"./2
F12880
Bolo 0
.40
c. .
w
o
w
118 0
.0/<
.4 .^ .08
8'-/0 0
o qm
O
24 0
.32
/, 2 ti .24
C. .
° 4v
0 v
b
.O6
/.6
g /6 u 60
u /2
V
2.0 .40
0
A20`40
11
1
.4 ^ .08
.2 .04
0
Airfoil: N,A.OA. 2309 R.N.:2,970,000
l/e%(R/sec.J: 71,0
Size: 5'x30"
Ires.(st7d atm.J:203 Dote:9-24-3/ -.2
Where tested.•L.M.A.L. Test. l!O.T.68/ -.4
Corrected for tunne%wa//effect
-8 -4 0 4 8 /216 20 24 28 32
Angle of attack, a (degrees)
0
Pim. 13,-N.A.C.A. 2909 airfoil.
12
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
,
St.. up ,.
/.25 2.24
2.5 3.//
50 4.3/
0
0
U
0
28 0
7.5
10
/5
20
25
30
40
50
60
70
80
90
5./B
5.86
6.89
].54
7.88
B.00
].77
7.14
6.21
5.02
3.62
200
95
1.09
20
L'w'r.
wo /o
0
-1.57
-2/6
-285
./2
m k _/^
o• x Test A
y
-326
-352
-3.82
-3.94
-3.99
-3.64
-3.45
-z.92
-231
-/.63
- .9/
0
20 40 60 80 /L
Per cent ofchord
7
l
5Z
C
/00 / i3 / 0 3
L L. Rod.: 1.58
Slope ofrod'us
0 /harddh2%/dr of
24 0 20
/.8 .36 ^.07
28 i
L6 .32 0.06
24
O
1.4 .26" : ^.05
l
20
.d
Q 20 Q 40
p 16, 60
l2 g 80
l2V .24 v.04
/6 p
40 .20 1)^.03
C
°
12
.8 w .16 0 .02
0 0
.6^ .12
.0/
6,0/00
.4'4 .08
Airfoil: N.A.CA.2312 R.N.:,1120,000
Size: S"x30"
Ve%(N./sec.): 69.2
Pres.(sf'nd.otm.):20.9 ofe:12-2-31
Where tested.' L.MA.L. Test.• V.D.T. 720
Corrected for tunnet-wall effect
-B
-8 -4
u
°
0
Ox
.2 .04 G
O 0 k -.2
a
4u
0Q
-4 0.C
'
32
VQ U
LID
k
36 v
.09
.44
2.0 .40 .V^-.OB
4 0
w
u
-.4
e
Airfoil: N.AC,A.N.:
23/2
R.
Dote: 12-2 3l
Test:
Corrected to infinite aspect
Q -.4
4 8 /2 /6 20 24 28 32
Angle of attack, a (degrees)
-4
O .2 .4 .6 .8 1.0 /.2
Lift co efficient,
F-B 19.-N.A.C.A. 2312.W.R.
Sla. Up'r. Lw^.
/.25 200 -1.,96
2.5 3B5 -2.74
50
7.5 6.28 -4.25
/ 0 7.06 -4.66
/5 8.25 -5.13
20 8.97 -538
25 9.3 -5.48
30 9.50 -5.50
40 9.22 -5.29
50 7.47 -4.77
60 7.36 -4.06
70 5.95 -3.22
60
90 1.36 -/.26
95 L30 ( .72
o
20
6-
o /o
Uv ,^.
0
lU
5.26 -3.66
6
O
2
O
40
60
60
P r cent ofchord
e
n
10
44
/00 /.161 (-./6)
/00 - O
v
L.E. Rod.: 2.47
5/ape afrodiu5
C
/.8 .36u.07
allC
.v. a
42V.24 •^ v.04
L/D
0
0
.4^ .08
u
4
Airfoil: N.A.C.A. 23/5 R.N.:3,06C108Ci
Size: 5"x30"
Ve1.(fl,/sec.): 69.8
Pres.(s%'d.oim): 2Q8 Date: 9-25-31
Where tested.• L.M.A.L. Test: Vb. T. 683 _
Corrected for tunne/-wo// effect .4
-4
V
-8
-4
0
4
6
/2
/6
20
24
Angle of allock. a (degrees) 28
32
0
u
0
0
40
W
•8 ^`
e
Airfoil: N.A.C.A. 2315
_.4
Dote: 3-25-31
'gCorrected
to infinite os
.4
/2 c
8e
.0/
.2 .04 Ok -. /
k
O
0 w -.2
0
U
C
-8
c
.8./60 .02
.6 u ./2
8 0 /00
0
/6 0
v
w
/.0U .20 u ".03
0 4v
o
28
V u
c.
l
t
1.4 .28-^ p.OS
24 k0 20
p /6
D6O
D o
°/2 60
36 v
l
32,
LL EE
i6
1.6 .32 p,06
2B 0 cho(dh^sot
' 20o40
.09
2.0 .40 G .OB
4.29 -228
O
U
48
:2
0
.2
.4
.6
.8
40
Lift coefficient, C
F1omz 16.-N.A.O.A, 2315 Wdoll.
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL
Sto. upr: C w1r.
0
0
1.25 /.// -0.60
Z. /.57 -/.04
5.0 2.28 -1.29
10 324 - 45
-B,
10
0
-1.44
-1.37
-1.25
- /.12
- .90
- .70
- .49
70 3.33
°
80 2.44 - 20
U
90 1.35 - .1 /
0
1061 (-.061
10
0.40
N
51ope ofnvdi,5
end of
2B p 0 through
chord: 2120
15 3.90
25 4.69
30 4.86
404.90
4.60
50
604.08
20
40
60
80
Per cent ofchord
/
00
N
01
44 .09
If
2.0 .40 06
U
v
/.8 .36 x.07
l.6 .32 0.06
C1 U
1.4 .26c th
24 20
I
I
20 v
/6 p
1.2,j.24.04
v
y -0Q40
13
t 0
4ap /0
20 4.37
/2
/.O,n .2011 12.03
C
LID
u
.8 m .16 0 .02
/6D 60
0 0
.6.".12
Q12-co 80
^.
8'.100
.4'.08
c
Airfoil: N.A.CA. 2406 R.N.:3, 120,000
Size: 5"x30"
3e/.(ft./sec.): 69.3
Pres. (sfnd. otm):20.7 Dote: 9-8 - 3/
Where tesled.• L.M.A.L. Test. V.D.T. 665
Corrected for tunnel-woll effect
C
-4 0.
'B U
-8
-4
0
4
8
12
16
20
24
Angle of ottock, a (degrees)
28
0
0
s
.2 .04 C-./
O
0 y -.2
° 48,
0Q
v
4u
.0/
-4 0
v
-8
0
u
.2
-.4-.4
_3r
,.rreC.A.
c4 c2 0 .2 .4
32
.6
.8
/.0
Lift coefficient.
/.2
/.4
FIGURE 16.-N.A.C.A. 2908 MEN].
5^. Up'r. LW
1.25 /.62 -123
2.5 227 -166
5.0 3.20 -2./5
7.5' 3.87 -2.44
/O 4.43 -2.60
15 525 -2.77
-,'b l0
Ug O
Y v -/0
O
25
6./8 -2.74
40 6.38 -2.62
30
-2.35
US
S.
RC 5.81
-2.79
50
605.
-2.02
704 27
-124
48
44
20 40 60 60 /G 7
Per cent ofchord
BO 3.10 - .85
90 1.72 - ,47
95 .94 - ,2B
O
/00 (./0) (-./0)
too - o
c0 G.E. Rod: 0.89
/
end of
28 p 0 Ih103gh
chord: 2/20
C
L6
C
.320.06
^q U
1,4 .28 0.05
N ^
3
L2,j.24M \.04
c. .
24 0 20
k
0 20 0 40
/6a 60
.09
2.0 .40 x.08
/8 .36 x.07
. 22 -/.63
O
i
40
44
L
.8 W .16
0
.6" .12
.0/
.4".08
0
Airfoil., NA.C.A. 2409 R.N.:3,1 /0,000
Size. 5"x30"
Ve).(ft./sec): 69.3
Pres.(sfnd.otm.):20.8 Dofe.3-9-31
0
-B
-4
0
4
8
0
/2 /6 20
0
24
28 32
Angle of ottock, a (degrees)
-4 0
0
.6 P
Q
0 m -. 2
Where tesfed:L.MA.L. Test.-VDT. 666
Corrected for tunnel-wall effect -.4
U
4s
U
.2 .04 OE-./
a
04
C
-4'4
w
.02
8'.100
'• C
° 4u
/2 c
0
/2g 80
-6
166
v
L0.20 ,03--
°u
v
9-3/
0 -.4
-4 :2
FIGURE 17, -N.A.C.A. 2109 Wrtoil.
.4 .6 .8 /.0
Lift coeff/cient,C
110,000
D.T. 666
14
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
So.
w,
!25 2.15 -465
2.5 4.13
299 -227
-3.01
50
7.5 4.96 -346
10 5.63 -3.75
15 6.61 -4.10
20 7.26 -423
25 7.67 -4.22
30 7.68 -4./2
40 7.60. -3.60
50 724 -3.34
60 6.36 -2.76
70 5.18 -2.14
60 3.75 -/.50
90 2.06 - .82
95 1.14 - .46
a
gi
31
-
V s° 0
vk-/O
p
0
20 40 60 80 /6 /
Per cent ofchord
d
48 .36 x.07
0
0 /6t 60
.8 u .16
.02
o O
.0/
.60 .12
.4 v .08
0
o /2.2 80
8 0 /00
k
° 4u
k
4 ^.
OL
°
.2 .04 V`-.l
C
.o
-6 U
16 8/2c
k
L
u
l
.c
?0 w
•-
/.2V.24^ v.04
v^
/.Ov.20 12.03
_4 0.
l
V U
1.4 .28- 0.05
24 a 20
k
V201 40
C 04
0
24 0
/.6 .32 o.06
'
L.E.Rad.: 450
v
Slope ofradius
Yh -d.- end of
28 p 0 chord:
2/20
l
36 ^
32
26 2f
.09
.44
2.0 .40 U .08
-4 0
0 k
v -.2
U
0
Airfoi/: N.A.C.A 2412 R.N.:.250,000
Size: 5"x30"
Ve/. (ft./sec.): 66.0
Pres.(stnd.otm.): 2/.0 Dote:/2-3-31 .2
Where tested.• &A.L. Test: V.O.T. 721 -4
Corrected for funnel-wall effect
-8 -4 0 4 8 12 16 20 24 28 32
Angle of ottock. a (degrees) •3
0-4
-4 -2 0
e
Airfoil:N.A.C.A
Date: 12-3-31
Corrected to
. .2 .4 `6
Lift cot
je
Test: V. 721 16
.sct ratio
10 /.2 1.4 16 /.8
^t. C
.
F-E 16.-N.A C.A. 2412 aic[oll.
S70o. Up'n GOY:
Z25 2.71 -206
25 3.7/ -286
5.0 5.07' -3.84
7.5 6.06 -4.47
6.83 -490
15 7.97 -542
10
I
20 8.70 -5.66
25 9.17 -5.70
30 9.38 -5.62
40 9.2S -5.25
50 6.57 -4.67
60 7.50 -3.90
70 6.10 -305
80 4.41 -2./5
90 2.45 - 417
95 1.34 - .66
100 (.16) (-.161
v
5/ope
5
too
-
^^ ^0
u t 0
O
20 40 60 80 /L
Per centofchord
g /s
o
1.6
C
ofradius
-8 u
o.06
c. .
LD
Airfoil: NA.CA. 2415 R.N.:3,060,000
Ve/(fL/sec.):698
Size: 5"x30"
Pres. (stud. otm.):20.8 Dote:9-10-31
Where tested'L.M.A.L. Test: VD.T.668
Corrected for tunnel-wolleffect.
-6 -4 0 4 6 12 16 20 24 28
Angle of ottock, a (degrees)
/6 0
/2
v
QUO)
u
c
.32,
^q V
0 4y
_4 0.
32
28 tl
1.4 .28-, 0.05
F 60
R 04
N
l
36 N
2.0 .40 .08
v
48 .36 x.07
o /2 8 0
80/0 0
o
40
.09
G
h rdh2%20 f
C 20' 40
44.
./I
44
L.E.Rod.. 2.48
28 0 l p
24 0 20
k
46
y kp _/0
.8 v .16 0 .OR-o a
.6 0 .12
.0/
k
0
.4 ^ .08
.2 .04 j-/
0 0 v -.2
°u
-.2
c -.3
-.4
0-4-32
.
-4 -2 0
Fmv .18.-N.A C.A. 2415 airfoil.
4^
u
°
0^
-4o
m
e
9-10-31
cted to
4 .6
Lift co
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL Si.. up'r. GWr
/0
r0
-2.45 ul U
-3.44
0-/O
-4.68
-5.48
0 20 40 60 80 /L 7
-6.03
-6.74
Per centofchord
20/0./5 -7.09
2510.65 -7/8
3010.68 -7./2
.44
.09
40/0.71 -6.7/
50 9.69 -599
D
60 8.65 -5.04
2.0 .40 x'•.08
70 7.02 -3.97
O
BO SOB -280
W
90 28/ -L53
/.8 .36 x.07
95 /.55 - .67
/00 /./9) (-./9)
/00 O
L6 .32 0.06
L. E. Rod: 356
v
ofrad'us
0
m
,
5"p,
'
end
28
0 chord: 2/20a t
1.4 .28`c 0.05
0
v
24 0 20
/.2V
.24kw^
m.04
a.P.
k
v
20 40
40v.208 L..03
L25
2.5
50
7.5
/0
/5
48
3.28
4.45
6.03
7./7
8.05
9.34
44
40
36 m
l
32 a
28 tf
.o
24 0
l
20 u
4
/6 0
4'.
G
N ^
0 6D60
a 1280
8'.100
n
Airfoil. NA.C.A. 2418 R.N.:3,060,000
Size: 5'x30"
Ve/(f1/sec): 69.8
Pres. (sfnd .1-):20.8 Date: 9-11-3/
Where tested.-L.MA.L Test: VDT. 669
Corrected for tunnel-wall effect.
-8
U
/z c.
u
L/D
04
0 Cv
c
-4 0.
u
15
k
.8v.16^1 .02-0 0
.6./2
.0/
k
.4 08
0
.2 .04 Of-./
0 0 w -.2
0
-.2
-.3
-.4
0 -.4
-8 -4 0 4 8 12 /6 20 24 28 32
-4
Angle of .Hock. a (degrees)
8.
4su
0^
`o
'4 0
v
_8
rn
%C KA.0 A. 2418
9-11-31
clad to inhi,it
V
.4 .6 .8
Lift coeff/cie/ C
n-RE 20.-N.A.C.A. 2418 airfoil.
sto. UP r. L'wr.
/.25 3.87 -2.82
2.5 S2/ -4.02
5.0 7.00 -SS/
7.5 8.29 -6.48
/0 9.28 -7./8
151
05
20 //.59 -6,52
/2/5 -8.67
3012
25
.36 -8.62
40
50
-7.31
60 9.79
19 -6.17
4 -0.67
70
80 74 1304
90 3.18 -1.66
as /.76 -1,06
/00(.
22)
/00 - (-.22)
0
L.E.Rad, 4.85
5/ape trod,, a
thro-1gh end of
chord: 2120
0
0
-6
U
l
tU
0
a
v
28 ° 0
3
Uc
U r 0
y o /0
4 a -/0
0
/0
.44
2.0 .40
v
L 2V.24
v
LO,v .2000
de
/D
h
0
.cv
k
U
0
° 4u
.2 .04
C 04
0 0
ti
0
.6x.160
.6'- ./2
.4.08
-4 0.
0
0l
/.4 .28
c.P.
o 12 yu° 80
8,0100
-8
a
ri
Ci
o /sa so
c
v
Q.
LB .36
/.6 .32
24 ,0 20
q20 0 40
20 40 60 80
Per cent ofchord
0
0
v
Airfoil.-NA.CA.242/ R.N.:3,000000
Size. 5"x30"
VeL (ft./sec.): 70.5
Pres. (sthd.Ol 7.) 0.6 Oate:9-//-3/ -.2
Where tested. 4.MA.L. Test: VDT. 670
Corrected for tunnel-wa/l effect. -.4
-8 -4 0 4 8 12 16 20 24 28 32
Angle of attack, Of (degrees)
Lift coefficient,C
FIOOEE 21. - N.A.C.A. 2421 airfoil.
16
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
.44
U
.40
0
t
u
0
0
.36
ac
.32
v
G
28 0
3
24
k
q 20F
.28-,
v
v .24.0
hW
v.20^
.0
U.12
o /2
v 80
.o
b,
v./6o
0 0
0 /6
08
.04
0
4u
OQ
C
-4 0.
u
-6
An of allock, a (degrees)
FIGURE 22.-N.A.C.A. 2608 airfoil.
5/0. Up'r. L'w'r.
0
0
25 1.57-1.27
2.5 Z2
224
5.0
3.75 -2.56
7.5 3-'0
/0 W. 2B -2.76
O
v
dfi
o
24
/5
20
25
30
40
50
60
70
80
90
95
100
D
O
20 40 60 80 /00
4
u
. /D
Pera- ofrhord
-2.97
-2.84
-2A4
-/.97
-/.50
-/.06
- .67
__
H
44
G$.
0
Airfoil:N.A.C.A.2509 R.N.:,R060,000
3e%(ft./sec.J: 69.8
Size: S"x30"
otm.J: 20.7 Dote: 9-1
Tres. (sfnd.
30"
teSMCd L.M.A. L. Test: U.D. T. 673
Corrected for funnel-wa//effect
Angle of oHOCk, a (degrees)
32
28
x.06
24:
p
got
/60
/2
Y.
8'^
4U
0j
°
-4 0
C"°/'
v
e
Airfoil:NA.CA. 2509 R.N.3.060,000 -/2
Dote: 9-/531
Test: l!O.T. 673 _16
Corrected to infinite aspect ratio
0 .2 .4 .6 .8 /.0 /.2
Lift coeff/cient,C
.2
-.4
-6 -4 0 4 8 12 16 20 24 26 32 °
/.4 .28 .05
/.2V.24u v.04
c
/.Ov.20, x.03
u
^
.8,) ./60 .02
.6 U .12^ .0/
0
.4 v .08
.2 .04 Vr -. /
/.6
e ndof
36
.09
2.0 .40 t'..OB
C
/.B .36 3.07
20
.c
-B
5.96
5/B
6.27
597
5.35
4.44
326
•/2
/O
.36
.99 - .22
(.10) (-./O)
/00 o
G. E. Rod.: 0.69
S/ope ofredius
throh
ug
0 chord:
2/25
Q0
Q 20040
0 /6^- 60
. 2 l2 80
9 o /00
° 4u
0Q
-4
SOS 2.9e
5.60-3. 02
Np
^$
.32
3 _3
E _' 4
c4 c2
FIGURE 23.-N.A.C.A. 2600 airfoil.
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE -DENSITY WIND TUNNEL
3.59
io
-/0
430
4.43
./
^ll.1LIJ_LJJ_J-L11J
0
ssz
20
40
60
80 /00
o v q
/
17
44
674
. l0
40
.44
.03
36 v
2.0
.40
• .OB
1.8
.36
1.6
.32
Per cent ofchorr(
y
4.44
%so
3.29
zs3
/.97
r
:42
0
e
of
24 R20
k
q20a 40
LID
8 .06
c.
.0.160
.02
Q)
l6 60
O
o
.6X ./2
.01
0
.4'.08
D
.2 .04
0 0
.2
Airfoil: N.A.C.A.25 /2 R.N.:3,080,000
O
VeG/flfsec.:69.4 _
_
Size: 5"x30"
Pres.(sfnd.ofm):21.0 Dote: /2-3-3/ ' 2v "?
Where tested: L.htA.L. Te5t. , VD.T. 722
Corrected for funnel-woll effect. -• 40 -•4
-.4 ..2
4 8 12 16 20 24 28 32
ng/e of oHock, a (degrees)
8 0 /OO
c
° 48
Op
C
-4 0.
-8 u
FrovaE 24. -N.A.C.A. 2612 airfoil.
l
O
U
0
O
C
N
26 0
3
24 0 2(
k
q
51a. Op".
0 - 0
/.25 2.63
-2.11
2.5 3.63 -2,94
5.0 4.96 -3.96
7.5 5.9/ -4.60
/0 6.66 -5.06
15 7.75 -5.63
20 8.48 -5.87
25
-5.92
30 6.92
9/9 -5.84
40 9./6 -535
50 6.62 -462
60 7.64 -3.76
70 6.28 -2.88
80 4.57 -/.9B
90 2.56 -/.07
95 L41 -.62
/oo
!-.lsl
/00 !./sl
0
L.E. RW..: 2.41
S/ape of, ailhro ugh end of
C& 2/25
u
U /0
C O _10
0
20 40 60 80 /L
NI#R
Per cent ofchord
UQ
1.4 .28
l
l
A
i
° o
.6X.12
k
.4'.08
LD
k 8 010(
0 4vC
C
-4 0.
ti
-B
.2 .04
C
u
l
0y
N
L 2V .24
m
LO ,y .20 0
c. p.
2004(
.o
.36
46 .32
o 6D&
0/28(
.44
2.0 .40
L0
0
0
Airfoi/.• N.A.CA.25/5 R.N:3,o60,000
Size: 5"x30"
Vet. fft/sec.): 69 8
Pres.(stndot. ):20.6 Doi-S-18-31
Where tested.-L.M.A.L. Test.-VDT 675
Corrected for tunnel-wall effect
-.2
-.4
-8 -4 0 4 8 12 16 20 24 28 32
Angle of attack, a (degrees)
FiomE 26.-N.A.C.A. 2616 airfoil.
270770-$ 6--3
l
20
160
a,
a
u
1
0 /2 86
24 0
Ci
1,4 .26^ 0.05
1.2V.241 m•04
0
LO x.200 Q.0.3
A
fins
0
p
k
C
a,
^-
32 $iU
28 2(
6.07
/31
a
26p
oa
33
I
v
12c
8lo
u
00
0^H. -4 0
v
8 eU
_/2
Airfoil: N.A.CA. 2512
R.N:3080,000
Test: V. D. T. 722 -16
Date: /2-3-3/
Corrected /o infinite aspect ratio
0 .2 4 .6 8 10 /.2 1.4 /.6 l.8
Liff coefficient, C
18
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
J/
.09
36
.08
32 aaP
.36 x.07
28 2f
.32 0.06
6
24 p
.44
U
ru
0
S
.40
v
v
V u
.26`c 0.05
•v a
Ci.24v N.04
28
24 lu
k
A 20o
20 m
16
v
p
u
0 /6^
ao
0.02
mo ./6 O
.
u ./2
.0/
.2 /2 u
k
4u
U
k
Bo
1
.08
0
00
.04
0 4u
-4 0
.o
0y
0
c
-4 0.
U
uu
o_
-B
__
.4 c2
FIGURE
36 v
VS
2.0
2B
7 5
3
7 6J
[. Rpd.: 4.es
JPe Ofrod'3S
o^d^'z% s f
.40
32 O
-.OB
LB .36 x.07
1.6 .32 x.06
^j
1.4 .28- 0.05
va
C
/.2V.24
C.P.
28
.6
24 0
l
20",v
168/2
y.04
k
_
4.03
U
11
.e^0 .lso0
i
.6'./2
.4 v .08
60.
0 4wu
x
.09
.44
s
28
24 0
k
420
Litt coefficient,
./G
112`
L
7 //
v
-16
0 .2 .4 .6 .8 /.0
2
0 20 40 60 BO /00
Per centafchord ^
U
e
-/2
Airfoil. N.A.C.A. 2518
Dote: 9-19 3/
Corrected to infinite -
2B .-N.A.C.A. 2618 airfoil.
mu ^o /0
0
0-/0
a
m
-6",
m -.2
m .3
-4
Angle of attack, a (degrees)
.2
CD
l
0
04
r
-4 0.
v
Airfoil. N.A.C.A.252/ R.N.:,^ /30,000
Size: 5"x30"
3el(ft/sec): 68.9 _
Pres(sfnd. ahn.): 2/.0 Date.9-2/-31 ' 2
Where tested.• LM.A.L. Test: l!O.T. 677
Corrected for tunnel-wa/l effect -.4
-8
-4
0
4
1-11.
/2
v .20 $ 4.03
6
/2 /6 20
24 28 32
Angle of attack, a (degrees)
.02
Bi
.0/
4^
k
0
0
OE -. /
0
-.2
.04
3
^
-4 0
v
e
.3
-.4
:4
10ote: 9-2/ 3/
I Corrected to infinite
0 .2 .4 .6 .8 1
Lift coefficient
1'..27.
-N.A.C.A. 2521 WdOU.
ratio
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE -DENSITY WIND TUNNEL 19
Slo.
0
[25
2.5
5.0
7.5
0
-
2.05
2.66
3.95
172
-1.73
-2.37
-a/7
-a66
10 5.34 -403
/5
6.25 -4.45
667
-4.6/
20
25 7.26
-462
7.51 -4.52
u U /0
Yo'
1U
4
0
p - /0
0 20 40 60 80 /C 7
Per cenfofchord
.44
30
40 7.56 -4.03
50 7.23 -3.36
60 6.56 -2.56
al
0u
kUO
2.0 .40
/
70 5.56 -/.77
80 4.15
90 2.34 -.56
95 1.26 - .33
/00
/.8 .36
/.6 .32
Ly
L.E. Rod.: /.56
v
V
-,of
28
3
24 1- 2(
k
q20o 4(
/.4 .281
v
1.2,j.24 ;O^
v
c'.9 2/30
choror
c. p.
LID
LO x.208
u
o.
sv.ls^
0 /sD s(
0
ti.
.4 08
8 104
4
0
.6X./2
o 12u 8(
C
0
04
.2 .04
O
h 04
Airf.A.., N.A.C.A.26t2 R.N.:3190,000
Size: 5"x30"
Ve4(ltlsec.):68.4 -.2
Pre s. (sthdotm):2/.0
Where tested: L.M. A.L. Test: VDT. 723
Corrected for tunnel-wo/l effect -.4
C
-q ti
V
-8
0
-8 -4 O 4 8 12 16 20 24 28 32
Angle of oitock, a (degrees)
Fxomz 28.-N.A.C.A. 2612 tWoll.
6t. UPY. L'^Y.
L25
204 -17S
L5 262 -240
6. 0 3.90 -323
7.5
10 4.66
5.26 -3.76
-4.12
20 6.73 -4.75
25 7./2 -4.77
736 -4.67
40
7.43 -3.47
50 7.13 -4.18
60
6.52
-2.6/
70
5.57
-1.67
804.42 - .83
/5 6.13 -456
v P /0
Ui O
h 0-10
0 20 40 60 60 /C 7
Per cent of
'd
.44
30
D
0l
l
2.0 .40
90 257 - .33
95 1.44 - .19
U
(-. 3)
L.ERod: /.58.
v
l ough end.(
26 o G 1h
chord: 2/35
O
(113)
1.8
.36
1.6
.32
O
100
1
Ci
1.4 .28c
.W
1.2, .24
/.o'.20
3
2402(
v
y
1
LID
c. .
0 Q 41
0 /6 6C
.8
8(
60/a
.4^ .08
0 4m
.o
.16
m
0 0
.6 ./2
.2 .04
u
0 p.
Airfoil: N.A.C.A. 2712 R.M-$060.000
Size: S"x30"
Vel.(ft./sec.): 69.5
0 Dote:12-4-31
P es.(sfnd.
c
-4 0.
ti
0
0
-2
Where tested.•L.M.A.L. Test: V.D.T. 724 -.4
Corrected for tunnel-wall effect
-8
-tl 4
0 4 8 M /6 20 24 28 32
Angle of attack, a (degrees)
M.. 29.-N.A.C.A. 2712 sftM.
Lift coefficient.(
20
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
Slo.
O
/.25
2.5
5.0
7.5
LPL,
-
L'wi-.
3.04
4.13
575
6.%
-/.07
-441
-L69
-/.66
cu
v o /0
U t 0
lU
p -/0
0
h
to Z90 -46/
15 9.15 -1.55
0
20 9.74 -1.74
25 9.92
20 40 60 80 /6
Per cent ofchord
-496
30 9.94 -2.06
t0
-/.55
U
k
v
28 i
0
0
24 a 20
k
20x40
A
.44
40 9.56 -2.05
50 875 -1.85
D
60
Z
70 6.13
-1.21
80 4.39 - .BS
90 2.4/ - .50
95 /.3/ - .3/
1.131 (-63)
/00
L.E. Rod.:756
5lope ofrodius
through end of
chor
2.0 .40
48 .36
1.6 .32
Ci
/.4 .,?8<
Li
d: -4//0
,N
2,j. 24 i?
v
c. p.
L/D
/0^.20u
I
I
l
U
6 16 60
U
.8m./60
0
.61" ./2
80
B 0/00
.4 '.08
M
0 4
o
u
2
0
0
.2 .04
0
4
0
0
Airfoil: N.A.CA.4212 R.N:3240,000
Siie: 5"x30"
Vel.(f//sec.): 68./
Pres.(sYnd.afm): 20.9 Dote:l2-/0-3/ -.2
Where tel...^L.M.A.L. Tezt:VD.T. 729
Corrected for tunnel-wo// effect -.4
-4. 0.
u
-B
O
-6 -4 0 4 8 12 16 20 24 28 32
Angle of attack, o: (degrees)
Lift coeff%c%eni, C
Hausa 30.-N.A.C.A. 4212 Udall.
0
0
-
20
A
Sta. Vp'r. L'w'r
/.25
12
Np
/,39 -.57
2.5 2.05 -.64
0
ti -/0
50 3./O -.54 p
7.5 393 -.35
/5 S. 30
7.00
.32
.68
B9
0
U
s5
.97
/6
I
I
C
OQ
c
-4 0.
-8
U
36
n,
08
.07
28 t
0
14:;z
l
20 m
OS
/. 2V.24^ti m .04-/6
v
1.0,t .20 4 .03
0
c.p.
40
.02
O
^
.6x .12
O/
.4 .08
0
k
n
p
/2'c
8.kC
0
4t
u
00
------
.2 .04 C
n
32D
0 06
U° u
1.4 .28, o
I
o 128 d
80/00
4U
cw
L8 .36
1.6 .32
40
09
2.0 .40 V
o /6 60
l
P
0
./0
44
/5
100 LO61 (-.O6)
/00
D
O
L.E.Rod.:040
W
51through
end
'80 0 chord: 4115L/Dof
24 ,0 20
0
20 40
80
60
Per cent of chord
%00
40 6.62 /.02
50 6.33 /.03
60 5.56
99
70 4.54 .86
BO 329
.65
90 /. ]9 34
D
q20
ffffH
0
/0 4.63 -.12
20 6.44
25 6.65
48
44
Airfoil. N.A.C.A.4306 R.N-W,90,000
Size: 5'X30"
VeL (f1./sec.): 70./ .2
Pres.(s £nd. of
3 Od1.:4-11-31
Where lested.• G.M.A.L. Test: V.O.T. 561
Corrected for funnel-wall effect _•4
-8 -4 0 4 8 12 /6 20 24 28 32
°u
c
F
0
S`
Angle of attack, a (degrees)
.3
-.4
' 4
-.4 -2
Q
Airfoil N.A.C.A. 4306
R.N.:3060.000 _/2
4-11-31
Test: V. D. T. 561 _/6
Corrected to infhde aspect ratio
0 .2 .4 .6 .6
/.0 /.2 /.4 /.6 1.8
Lift coefficienl,C
-
40
w
0 0 v
Fluvaa 31. N.A.C.A. UN aidoll.
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL
cows
0
44
20
40
60
60 /00
(-.09)
c
:T8_9
28 0
s
C
rodi3s
f
42 V .24
24 0
t
28
O
24 p
t
20 u
CL
v.04
16p
C
n
wo qvu
00.
Airfoil: N.A.CA. 4309 R.N.:3,060,000
C
Size: 5'X30"
Ve%(H./sec.): 69.8
Pres(stnd. otm.): 20.B D&e.4-/3-31
Where tested.• L.M. A.L. Test.• VD..T 563
Corrected for tunnel-wall effect
-4 0.
-8
0
8
4
/2
/6
20
4yU
01)
.4 .08
0
.2 .04 J-.
Ci - /
0 O
2
u
-2
' 2^ ' 3
ff_4
-4
n
o
24 26 32
..4
Angle of o{tock, a (degrees)
c ^° -4 0
W
B
Airfoi/:N.A.CA. 4309
R.N.:3060,000 -/2
Dote: 4-133/
Test: V. D. T. 563 _/6
Corrected to infinite aspect ra{io
-
-2
Fioun 32.-N.A.C.A. 4308 Wdoll.
51a. Up'r. L'wY.
O
- 0
L25 2.64 -1.29
2.5 3.63 -/. 75
5.0
7.5 S6.22
/0 7/2
/5 846
20 9.34
9.62
25 /0.00
3o
N p
0
U ^
0 _/o
-2.19
-2.34
-2.39
-231
-2.17
-2.07
-2.00
0
20 40 60 80 /L )
Per cent fchord
.44
70 6.39 .95
60 462
- .64
90 254 - .38
95 1.38 - .25
(' 13) (-Q3)
100
L. E. Rod.: 1.58
m
C
2.0 .40
/8 .36
/.6 .32
si oe orroe^s
through end of
chord: 4//S
26p;
O
U c/0
40 9.75 -1.86
50 8.98 -/.6/
60 765 -1.28
-
L
l
O
l
U
k
O
w
2f
Ci
1.4 .28 c
N
LID
/.2V .24
C.P.
v
LOv.20^
q20p41
u
/6' 61
t
0
0
o /2 ^ 81
60 /Ol
0 4m
.o
u
v
04
C
-4 R
-8
8t
.6 ku ./2^ .0/
p /2 U
24
32 O
Co,
4O v.90^ .03
u
.8 i .16^ 02
p /6U
a
36 oip
Q) 0
0200
t
^
40
44
09
$
2.0 40 1.08
w
).8 .36 u.07
w
1.6 .32 0.06
cj U
/.4 .28 c 0.05
ao
U
.l0
Per cenfofchard
- .05
`o
48
o -l0
- %J7
-1.28
- 1.0/
- so
60
-_
-- 29
1s
- Os
D
./2
o i0
0
- 34
21
.4^ .08
.2 .04
a
0
Airfoil: N.A.C.A. 4312 R.N.:3,/80.000
Size: 5"x30"
Ve/.(fl./sec.: 68.7
Pres.(s{i+d.otm):20.8 Dole:12-14-31 -.2
Where tested: L.MA.L. Test: V D.T. 731 -.4
Corrected for tune/-wall effect
0
-8 -4 0 4 8 12 /6 20 24 . 28 32
Angle of o{lack, a (degrees)
FxoII z 33.-N.A.C.A. 4312 aWoll.
0
.2
.4
.6
.8
1.0
Lift coefficient.
1.2
/.4
1.6
18
tz
22
REPORT NATIONAL ADVISORY COMMPPTEE FOR AERONAUTICS
sto. UPS'. 4M,
/.25 3.32 - 60
2.5 4.47 -2.26
5.0 6./3 -2.97
7.5 737 -3.3/
D
v
0
73/
-L66
0
100
-4
V
l
V u
20
va
c.
1.2 Ci.24X j.04
p.
v
m o
L/D
/. 0,E .20 4.03
u
.8 v./600 .02
`c
o 'o
s °./2
k
.0/
.4".08
0
.2 .04 G-./
0 0 k0 -.2
-.2
w-.3
v
Z-4
-.4
0
m
-8
l
32
28 i?
.d
24,
44 .28`c 00 .05
chord:4 /5
420040
N ^
e /6 60
/2^8 0
B o /00
e
4
36
1.6 .32 0.06
C
24 0 20
4
CR
40
.44
.09
S
2.0 .40 G.08
,v
1.8 .36 x.07
L. E. Rad.: 2.48
Slope afradius
0 through end of
0
0
t0
BO 529 -/.30
9 0 292 - .74
95 1.5B - .44
100 (.l6) (-.16)
D
z°
20
40
60
BO
Pe Cent of,or,
50/0. / -2.93
60 9.01 -2.42
70
44
/I
20 /0.60 -a60
25 11,32-3.55
-3.50
301/.
4011.18 -3.33
L
-c°
k
W
48
0
/0 0.36 -3.50
r
l5 9.85 -3.62
28
t0
v uko -/0
y
Airfoil., N.A.CA. 4315 R.1i:3,120,000
Size: 5"x30"
Vel.(ft./sec.): 69.3
Pres.(stnd. aim): 20.8 Dole:4-14-31
Where Msled.•) A.L. Test.-VDT 565
Corrected for funnel-wall effect
-6 -4 0 4 8 12 16 20 24 28 32
Angle of otfock, a (degrees)
4^
00
-4o
v
-8
Airfoil: N.A.CA. 4315
Lift coefficient, Q
FIOREE 39. -N.A.C.A. 9315 Elrtoll.
Sto. Up 'r. L0,
0
-
0
/.25 4.07 -/.90
2.5 S.36-2.74
S. 7.20 -3.68
Z58.56-4.25
W b0
t 0
y -/0
0
0
/0 9.64 -4.56
/5 //.22 -4.92
20 /225
2512.
80 -500
30 13.00
40 /264 -477
20 40 60 W
Per centofchord
.44
50 //.65 -4.24
60 10.15 -3.55
70 6.24 -2.76
D
0
kO
a
)
/0
v
60 5.95 -/.94
90
329
-/.OB
48 36
95 1.79 - .64
100
(-./9)
too (./9)
- o
v
28 p 0
3 20
24
q 200040
N ^
o /6' 60
0 /2 oCu 80
/.6 .32 0
L. C. Nod.: 356
Slope ofrod/us
Through end of
chord: 4 /5
L;,
Ci u
1.4 .28 o
va
1.2V.24w m
c.p.
40.^.20va 0
u a4
LD
0
.o
u
4 04
C
4 0.
-B u
0
.6' .12
.4'.08.
8 0100
0 4v
V$
2.0 .40
.2 .04 V`
C
0 0 y
Airfoil: N.A. C.A. 4318 R.N:3,09Q000
Si- 5'k30"
VeZ(ft/sec.): 69.5
Pres.(st^7datm.):20.8 Dofe:4-14-31 -.2
Where tested.-L.MA.L. 7-est.-VDT 566
Corrected for tunnel-wall effect -4
U
v
0
-8 -4 0 4 8 12 16 20 24 28 32
Angle of attack, a (degrees)
FtomE
Lift coefficient, C
35 -N.A.C.A. 4318 MO.
23
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENBFPY WIND TUNNEL
Sto:
0
o
1.25 4.84
-2./9
25 129 -3./8
50 6.26 - 4.38
,0/0. - 564
15 1262 -6.19
2013.72
ut / 0
l
o -/0
R
0
-42
S.
25 /4.29
1450
30 /4.09
40
50 12.96
60//. /
70 9./8
D
./2
44
20 40 60 80 /C 7
Per cen/ofchord
-6.50
-6.50
-6.23
-5.56
-4.66
-3.66
80 6.64 -258
90 3.67 -1.43
95 2.00 - .84
t
k
O
N
C
o
c.p.
LID
/2 °C80
u
vi
8'.100
C
0 4w
u
°
1.4 .28-, 0.05
20 w
1.2V .24 x.04
a '^
/.0- .20 u x.03
u
a
.8 v '16° .02
o O
.0/
.6X./2
k
0
.4".08
/6 0
a
W ^
/6O 60
M
28 b
C°
ihrou9 =
28 0 0 chord:
4115
24020
20 40
32 O
k
1.6 .32 0.06
o
L.E. Rod. 4. 65
s/ape '
1 ,
v
36 m
N
18 .36
/00 (.221 (-.22)
/00 -
D
h
.44 .09
a
2.0 .40 V .08
4y
0
O'er
0
Y
2 .04 G f
C
^° O 4
C
0
c -.3
Airfoik NA.CA. 4321 R.N:3/20,000
Size. 5"x30"
Ve%(f/./sec.): 694
Pres.(sfnd.ofm.):20.7 Dote:4-15-W -.2
Where lested.-L.MA.L. Test: V.OT 567
4
Corrected for funnel-wo/1 effect
-4 0.
8 u
-4 0
0 m-.2
0
1-4
-4
-8 -4 0 4 8 12 16 20 24 28 32
Angle of oltock, a (degrees)
- 81
R.N. 3,120, 0001
A/rfo/1.NA.CA 4321
Test, VD.7.5671
Dole:4-15-3/
^ Corrected to infinite ospectrobo
0 .2 4 .6 8 /.0 /2 1.4 16 LB
L/ft coeff/'c/enl, C
.
Ftovaa 38.-N.A C.A. 9321 SIM%
D
u
k
O
.
sf. Up', L''wr
D /0
JS 1.25 -.64
2.5 1.68 - 79
u u 0
5.0
i0
/5
20
25
30
40
so
60
70
2.79 - .82
44
ass - .60
O 20 40 60 80 /G
5.15 -.25
Per cent ofchord
5.90
6.42
6.76
6.90
6.55
5.95
485
.12
.46
.74
1./0
1.24
1.27
1.16
1.96
.49
80 3.56
90
48
0
u,A
40
N
95 1.05 24
/00 (.06) (- 06)
o
00 L.
E. od.: 0.40
w
51ape ofrodi35
through end of
2B o ( chord:
4120
1.8
V
W
/.2V.24 v.04
m
lo-.20u x.03
L/D
c.p.
168a
12'^
u
.8 ^./6 0 .02
t
.6 0 ./2 .0/
k
.4".08
0
.2 .04 0-./
O O LIZ -.2
u
80/a
`^
20'.
1.4 .28- 0.05
O 2004(
/6 6(
^
28 ti
.d
246
l
.36 ^.07
1.6 ,32 0.06
C
24a 2(
36 w
32
.44
.09
a
2.0 .40 U.08
C
I
in
4u
^ 0 Qa
c
°
,o
C,
Airfoil.-N.A.CA.4406 R.N:3,10Q000
54, 5"x30"
Vef.(ft/sec.): 69.4
Pres.(sfd.
nofm.):20.9 Dote:8-21-31
Where tested: L.MA.L. Test: VOT. 651 -.4
Corrected for tunnel-wa//effect
-4 4
U
-B
8
-4
0
4
B
12
16
20
24
Angle of oltock, a (degrees)
28
32
F[0vae 37.
4s
u
0Z
O
-4o
v
e
U
0-.4
-4
-N.A.C.A. 4908 airfo.
N.A.CA. 4406
-21-31
L ift coeff/'c%ni. 0,
R.N..:3,/00,000
Test: VO.T. 651
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
24
Sto. up-,. LWr.
0
/.z5
-
O
/.e/ -LOS
2.5 2.61 -1.37
S. 0 3.74 - /.65
75 4.64 -1.74
/0 5.37 -/.73
a cu
10
u o
48
2
0
R o-/0
0
/5 6.52 -155
44
20
40
60
BO
/
0
Per centofchord
20 7.33 - /.30
0
40
09
36
25 7190 -L02
30 8.25 - .76
40.8.35- .35
44
50 767 - .07
60 7.00
.14
70 5.76 .26
D
2.0 .40
60 4.21
.26
90 2.33
.14
95 1.26
03
100 (.03) 1-.091
0
a
/00
v
28 'b 0
0
L. E. Rod.: 0.69
Slope o(rodius
through end of
chord: 4 20
32.
1.8 .36 .^ .07
28
1.6 .32 0 .06
24
V u
1.4 .28 c o .05
20
V ^ .614
16
W a
3
24 a 20
020040
.08--eat
c.p.
Lm
LO,v.20u
y ^
.03
U
.$ v./s o
8.
r°
p 16 60
o /2.g 8 0
.6 U.12
0/
8 0/00
.4'.08
0
0
0 0 m -.2
-4
-8
0
.02
0
w
c
0° 4 u
04
C
4'
.2 .04 t
c
0
AirfOlL'NA.CA.4403 R.N:3,170,000
Size: 5"x30"
Ve1.(ft./sec.): 68.6
-4 0.
Pros.(stndatirz):'WA2/.0 Dote:8-24-31 2
V
Where tesfed L. .L. Test: V DT 652
-6
Corrected for tunnel-watt effect -.4
-8 -4 0 4 8 12 16 20 24 28 32
Angle of attack, a (degrees)
c
c
.3
_ 4
-4 -2
Airfoil. NA.C.A. 4409
R.N.:3,170,000 12
Test: V. D. T. 652 -/6
Dote: 8-24-31
Corrected to inf'nite aspect ratio
0 .2 .4 .6 .8 /.0 /.2 1.4 16 /.B
Lift coeff/cteni,C
Fmm. 38. -N.A.C.A. 4408 I& M.
S1o. Opi-. Lm,
ac
o 2.44
-0
1.25
-/.43
2.5 3.39 -1.95
5.0 4.73 -2.49
7.5 5.76 -2.74
^r / 0
h
y -/0
0
0
/0 6.59 -2.66
/5 7.69 -2.66
0
ku
a
mc
20
40
60
60
/0 0
Per centofchord
20 6.60 -2.74
25 9.41 -250
30 9.76 -226
40 .460 -/.80
50 919 -1.40
60 8.14 -L00
70 669 -.65
80 4.89 - .39
.44
2.0 .40
so 271 - .zz
95 /.47 - ./6
/.8 .36
C
l00 1.131 (-./31
100 -
O
1.6 .32
L.E. Rod.: 1.56
5/ape of-d,11,
V
endpf
2810 0 lhrough
chord: 4 20
3
24 1- 20
LID
w
Q20 a4 0
l
ri 16 6 0
12 8 0
80/00
1.4 .28
W
A2V.24
c.p.
v
LO y.20°
U
$ .!6
0
.6'./2
.4'.08
`^ C
° 4u
.0
R° 0 4
C
_4 0.
_g v
u
0
.2 .04
C
0
0
Airfoil, N.A.CA. 4412 R.N:3,200,000
Vel.(ft./sec): 66.6
Size: 5"x30"
Pres.(sfnd.oim.):20.8 Doie:12-15-31
Where tested: L.h1A.L. Test: V D.T. 732
Corrected for tunnel-wail effect
-.2
-.4
-8 -4 0 4 8 12 16 20 24 28 32
Angle of attack, or (degrees)
-
1NOME 30 .-N.A.C.A. 4412 91doil.
Lift coefficient. 0.
3
3
a
i
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL
S0. UP r. L'0r.
/.ZS 3.07 -1.79
2.5 4./7 -2.48
5.0 5.74 -3.27
75 6.91 -3.7/
/0 7.84 -3.98
15 9.27 -4./8
^D l0
U
lU
4p
O
0
C
28
44
20 40 60 BO /0
Per cent ofchard
.44
5 /0.53 -2.72
60 9.30 -214
70 7.63 -/.55
90 308 - .57
95 1.67 - .36
00 (.16) (-.16)
L. C. Rad.: 2.48
slope o7u9
through end o7
Chord: 4120
iW 0
.
24 '3a 20
q20040
0 16--60
.09
28
/.6 .32 0 .06
V u
1.4 .28 o^.0$
W ^
/.2 V . 24k m 04
24'
Lc
e.p.
ZT
.03
A./6^
.02
L/D
/.0.200
0
oe
.6u .l2
.4'.08
0
0
0 0 m -.2
-8
2
cv
0
5/a. Up •r. Lm,
° 376
- -2//
°
/.25
2.5 Soo -299
5.0 6.75 -4.06
7.5 6.06 -4.67
/0 9.// -506
15 10.66 -5.49
20 //.72 -556
25 12.40 -5.49
30 12.76 -5.26
40 12.70 -4.70
s0 /1.65 -402
60 10.44 -24
3.
70 6.55 -2,45
80 6.22 -/.67
90 346 - .93
95 /.89 - .55
a
O
k
.6
.B
LO
L2
/.4
L6
0
/.8
48
20
40
60
80
Per centofchord
/0
0
40
/0
.09
h
36 w
2.0 .40 ^ 08
AS .36 ^.07
28 Z{
1.6 .32
24 0
44
a
32 v
N
C
g
0.06
G u
/.4 .08- 0.05
If
l
20 m
• v .04
1.2V .24 F
v `^
40.y .200 ^.03
168
a
v
2.'c
k
.0
.8v.16^ 020 0
6 u .12
.0/
k
0
.4 08
`•
. 2 .04
n
Angle of attack, a (degrees)
Oa
C `
0 0 m -.2
0
u-.3
v
0 -.4
-8 -4 0 4 8 12 16 20 24 28 32
4^
3
Airfoil. N.A.CA. 4418 R.N.:3,100.000
ve/(f//sec.): 69.5 -.2
Size: 5"x30"
Dcle:8-27-3/
I-rl'
Where tested: L.MA.L. Test: l!0.T 655 -.4
Corrected for tunnel-wall effect
.4
FIGURE 41.-N.A.C.A. 4418 airfoil.
270770-35-4
.4
Lift coeff can't C
44
L/D
u
.2
W k _/0
Co
c. .
0 4 Cv
720
-N. A.C.A. 4415 airfoil.
uso l00
4.E.Rad.:3.56
Slope ofrod)u5
Z'
D through endof
chord
28 0
24 0 2 0
q 2004 0
N ^
/6^ 6 0
0
o /2r8
u
8 0/0 0
R.M.: 3,110.000 /2
Airfoil. NA.CA . 4415
Dote:8-26-31
Test: l!D.T. 654 -16
' 4Corrected to in fihite aspect rofo
_
-4
FIGURE 40.
U
8'
4'
.0/
-8 -4 0 4 8 • 12 16 20 24 28 32
Angle of attack, a (degrees)
-8
2
a
0
P(sti7d.otm):20.7 Dote:8-26-31
Zr. tested: L.MA.L. Test: VOT. 654
Whe
Corrected for tunnel-wall effect _4
-8 u
c
l6
0
Airfoil: NA.C.A.44/5 R.N:3 /10.000
Si- S"x30"
Vel.(fl/sec.): 69.5
.c
44
04
20
.2 .04 c
.o u
a0 0 4
4
32.
_6
LB .36 ._6 .07
`•
0 4Cv
W
36
°°
2.0 .40 U .08
0
x 80/00
2
40
./0
80 5.55 -/.03
k
O
48
2
30 //.25 -3.75
//.25 -3.25
U
25
0
-l0
20 /0.25 -415 25 /0.92 -3.9B
D
-4 o
-B
-/2
I: NA.CA. 4418
R.N.: 3,/00,000
Test: VD. T. 655 -16
9-27-31
cted to infinite aspect ratio
.4 .6 .8 LO L2 1.4 /.6 /.8
Lift coefficient, C
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
26
D
50
BO
70
60
90
w _/0
hW p
t
y
-36 v
.09--
8
2.0 .40V .08
u.07
26 e
0.06
24 0^
/4 .28'^ 005
v ^
20 u
Q
1.6 .32
O
V U
C
3
241,20
^
32 U
a;
N
1.8 .36
100 /.- F0
l0 0
0
LG Bod.: 4.65
v
Slope ofradius
end of
28 0 through
chord: 4/20
/.2V.24 y v.04
/.'O w .20 um °.03
R
U
B v . /s 0 .02
o
°
.6 u ./2
.0/
S
0
.4".08
.
r-
c. p
L/D
b
40
r
n
`o
o /6^ 60
./0
.44
-4.40
-•235
-23/
-/.27
Q 20x40
y ^
48
44
0 20 40 60 80 l00
Per cent ofchord
-6.75
-6.9B.
-6.92
-s.76
-6./6
-5.34
13. lB
//.60
9.50
6.9/
3B5
./2
10
g0
D
]5 924 -5.62
/B %35 -6.15
12.04
15
2013.17
1314.27
B6
25
30
4014.16
-
zo
L^ r
-2.42
125 4.45
2.5 5.B4 -3.46
50 7.82 -4.78
0/280
8 0/00
168
12
^
0
4
uk
0
0
z
0° 4 u
0Q
C
C
0
-4 0.
-B
-8
0
-.2
Airfoil NA.CA. 442/ R.11:3.110,000
u
Size: 5"x30"
Ve/(H/sec.): 69.4
Pres.(stird ofm): 20B Dote:8-28-31 .2
w ' 3
Where tested: L.M. A.L. Test: VD.T 656 -.4•4
Corrected for tunnel-wall effect
R- -4
-4 0 4 8 12 16 20 24 28 32 72 0 .2
Angle of attack, a (degrees)
FIGURE 42.-N.A.C.A. 4421
-8 u'
e
R.N:3,110,000 /2
Dote. B-28-31
Test: VD.T. 656 -16
Corrected to infinite aspectrotio
Air foil, N. A. A. 4421
.4
.6
.8
LO
/.2
Lift coefficlent,C
1.4
airfoil.
5f. UPr.G' r. 10
D
/. 25
-. 7/
'
/0
362
-.89
/5 4.74 -.64
20545
-:32
25 59B
.34
30
6.36 .02
40 6.74 ,93
50 6.65 /.35
60
613 /.56
70 5.21 /.53
60
390 /.25
90 2 /8 .72
95 1.17
35
/- /-0sl
i°o°o
L.E.
Rod.:
0.40
51ope od
fro us
UC 0
.66U _/0
5.0 260 -/.00
hp
2.5
7.5 3.25 - .97
0
U
W
e
0 20 40 60 80
Per centofchord
24 0 20
k
Q 20F40
V ^
C
c. p.
c
-4 0.
_8 d
'01N
/.2V.24^ y.04
/60
k
0.03
uo
u
.81./60 .02
.6 u .12^
w
8,0/00
k c
20 4 u
0Q
32 O
28
24:9
12'
/.0.1.200u
12- BO
.o
36-0
ro0 0^
y
L D
g` /6't 60
a
40
.44
.09
2.0 .40 -08
1.8 .36
.07
/.6 .32^ x.06
44 .28 , 0.05
28 p 0 through entl of
chord: 4125
44
00
,o
.0/
.4 .08
4u
0
0
r
°
0
O
Airfoil: NA.C.A. 4506 R.N,:$05Q000
Size: S"x30"
Ve%(H./sec.) 70.0
d o tm.): 20.6 Dafe: 4-/5-31 •2
Pres.(st'n:
Where fesfed• L.A.L.
M. Test: V.O.T.568
Corrected for tunne/-/l effect -.4
v
O y -.2
°u
c •3
- 4
-8 ^`
e
Airfo#-N..A.C.A. 4506
R.N:3.050,000 -l2
Oote: 4-/5-3/
Test: V AT 568 _l6
Corrected to infinite aspect ratio
-8 -4 0 4 8 /2 16 20 24 28 32-4 -2 0 .2 .4 -6 .8 /.0 /.2 l.4 46 /.8
Angle of otfock, o: (degrees)
Lifl coefficient.
FIGURE 43. N.A.C.A. 4803airfoil.
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL
up',
13S
2.47
3.54
4.36
SOS
6.12
6.89
51.
/.25
2.5
5.0
7.5
/0
/5
20
V0,
-1.12
-1.50
-1.84
-1.99
-2.05
-/.96
-1.75
D /O
U ^ 0
L, -/0
20 40 60 80 /0
Per cent ofcho^d
0
25 7.46 -/.47
30
.52
,40
6./9
50 7.97
0
80 4.56 .60
90 256 .35
95 /.
/5
(.09) f-.091
/00 0
L. E. Rod.: 0.69
Slope afrad'us
through end of
.03
.42
.62
60 7.27
70 6./2
U
k
.44
7.85 -1.16
-
a
2.0 .40
/.e .36
100
v
28O
3
24 0 2i
/.6 .32
C
V
/.4 .28c
chord:4 25
,W
/.2V .24
LID
w
N 20041
v
LOv.20^
p /6 61
8x./60
D
.6h ./2
.4'.08
o O
o
i
°12^Bi
8 0 /Oi
c
C
.2 .04
° Su
o
v
^ 0 Q
O 0
Airfoil, N.A.C.A. 4509 R.N..:3,120.000
V.I.(fG/sec.): 69./
Size: 5"x30"
P es.(s fnd. almf:20.9 Dote: 4-15-31 -.2
Where fested: L.MA.L. Test: V D.T. 569
Corrected for tunnel-walleffect -.4
.0
-4 0.
G
-B
-8
-4
0
4
8
12
16
20
24
Angle of attack, a (degrees)
28
32
Lift coefffcient,Q
Proves 44.-N.A.C.A. 4608 airfoil.
5o. Up', L'-r.
125 2.33
25
SO
7.5
/0
3.22
4.50
5.46
6.25
vD /0
O
-2.07
-2.67
-2.99
-3.18
O
0
15 7.46 -3.28
20 a34 -3./7
30
40 9.64 -/
-/.97
50 9.29
.29
606.43 - .70
20 40 60 60 10
Per cont ofchord
-2.95
25 8.95
9.37 -266
Ul
.44
2.0 .40
70 7.06 -.2,9
805.23 -:04
90 2.93 .00
95 /.58 - .04
/00 (./2) (-. 12)
/00 - 0
LX. Rod: 1.56
Slope ofrodus
through a 2Z
°
t
U
aC
v
28 0
48 .36
1.6 .32
C
Ci
/.4 .281
chord: 4/25
3
24 L. 2
W
2,j. 24 i?
C.P.
LID
^o
/.0 r .20 0
X20041
N ^
u
.6 .160
0 0
o /6^ 61
o /2 C
u
e
r
.6°./2
k
4".08
B o /Oi
0 4v
.2 .04
u
o
v
^ O4
C
AlrfOlL'NA.CA.45/2 R.N:320Q000.
Size: 5'k30"
Ve1.(fL/sec): 68.2
Pres.(sfnd ofm):2/./ Dafe:/2-/6-3/
Where tesfed'LAIA.L. Test:V.DT 733
Corrected for tunnel-wall effect
4 0.
v
-8
- 0 -4
O
4
8
12
/6
20
24
Angle of attack a (degrees)
0
0
-.2
-.4
28 32
Lift coefficient,C
F'lovac 46.-N.A.C.A. 4512 airfoil.
27
28
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
St.. Up". L'w^-.
0 - o
/.25 2.94 -/.88B
2. 5
00 -2464
0 4
5. 49 -3 8
7.5 6.60 -398
4
to
20
wuo ^/0o
].4s -4.30
IS 8.85 -4.59
20
-450
0
2B
ti
L6 .32 x.06
24 p
/.4 .28- 0.05
20 u
va
1.2V .24^ v.04
v
/.O.w.20' x.03
U
1 h
.B m'/6^ .02
o '0
.6 wu .12 .0/
O
.4J .08
Y
LID
/6^ 60
o /2c80
8,0100
/6g
v
12
6 ^^
k
4
0
-4 0
v
O 0 w -.2
04
C
Airfoi/-NA.CA.45/5 R.N.:3,080,000
Size: 5"x30"
Vel tft./sec.): 69.6
P es.(sfndofm.):20.7 Dote:4-/6-3/
Where tested'L.MA.L. Test: VD.T. 571
Corrected for funnel-wall effect -.4
-4 R
v
-8
J1
4y
.2 .04 V-./
C
u
28 ZS
.d
C.P.
0 4 Cv
a
32 O
v
o
24 0 20
.o
2.0 .40 •x.08
Ci
100 O
L. E. Rod.: 246
Slope o/ radius
I=
9,5
end
0 chord: 4125of
l
40
N
36 w
ti
v
48 .36 x.07
-1.e5
70 299 -1.20
BO SB9 - .69
90 332 - .35
y 20040
.09
.44
60 9.56
s°
cN
7
20 40 60 BO 10
Per cen/ofchord
3010.80 -4.16
4011.019 -3.44
501062 -262
D
48
44
9.6/
25 /0.47
O
.12
0. o
-B
U
0-.4
-8 -4 0 4 8 12 /6 20 24 28 32
Ang/e of Mock, a (degrees)
-4 -20 .2 .4 .6 .8 /.0
Lift coefficient,C
y
Fca sE 46.-N.A.C.A. 4518 airfoil.
Slo. UP 'r. LWr
0
0
/.25
25
3.56
4.79
6.49
50
/0 6.74
/5/0.25
20 //.27
30 12.37
40 /253
50 1/.94
60 /0.72
70 6.92
60 657
-2.25
-316
-4427
-5.4/
-569
-6.01
-5.
9
-5.66
-4.69
-3.94
-296
311
-1.34
90 3 69 - .7/
95 2.01 - .44
/00 (.19) (-./9)
/00 - 0
L. E. R.d.d.: 3.56
Slope of-d,,,
2511.96
s
k
a
v
26 0 0 c hord
24 ,b 20
O 20 0 40
y ^
/6^ 60
a cu
/0
y a
4
l 0
00
0
.//
I
20 40 60 80 IC
Per centafchord
44
2.0 .40
C
0 4v
o
u
04
c
-4 0.
-8 u
.09
36 w
l
- -
.x.08--_.._
v
..._.......
Cno
/B .36 y.07
k
1.6
4125of
168
L2^.24X v.04
v
1.0 .a .20 4.03
U
.8v./6o 02
0
.6 .12 0 .0/
c.p.
IL
32 a
28 ti
.32 0.06
C
1.4 .28`c 0.05
W a
o /280
60/o0
`^
40
''
.4 .08
k
6i
a'
4y
u
4
0
0
.2 .04 V`-./
0 0 m -.2
Airfoil N.A.CA. 45/6 R.N:3130.000
0
• Size: 5"x30"
3e).(ff./sec.): 66.9 .2
.3
Pres. (sthd afm): 2L0 Dote:4
Wherefested.-L.MA.L. Test.VOT 572 _
_.4
.4
40
C
Corrected for tunnel-wc , H effect
-8 -4 0 4 8 12 16 20 24 28 32
Angle of otfoch, a (degrees) m
e
Airfoi/: NA.CA. 4518
Dote: 4-16-31
R.N: 3 /30.000
Test, VD. T. 572
Corrected fo inflhVie ospectrofio
-4 -2 0 .2 .4 .6 .8 /.0 L2 14 /.6 L
Frooxx 47 .-N.A.C.A. 4518 airfoil.
Lift coefficient, 0
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE -DENSITY WIND TUNNEL r
r.
sl. UP' . L'v,
1.25 4.23
2S 5.60
SO 750
7.5 0.90
/0.00
/0
1511.63
20 1273
S
0
m
26 O
0
24 t. 20
k
Q20E
y o 40
-2.56
-3.B6
-505
-5.90
-6.50
-7.17
-7.42
2513.46 - 7.38
30'3.06 - Z/6
40 1397 -635
60 13.27 -5.27
50
11.86 /2
70 9.65 -4.
-3.00
80 7.26 -1.97
90 406 - /.05
95 2.24 - .63
/00 (22) (-22)
100 - 0
LE.Rod.: 4.85
5/opeofrodius
through end of
chord: 4/25
/2
94Mm
N kp -/O
0
0
311
.08
W
L8 .36 o .07
h
32 D
28 tt
1.4 .2Bc 0 .05-
0 0 y
60
7.44
28
0
_4
3
c.p.
-8 u
-8
-4
.2
.4
.6
.8
LO
Lift coeff/cient,0
1.2
/.4
t.6
L8
toll.
X140
CN
/.8 .36 .0
0
+-H 24
1.4 .28 C 0
N
L2V .24w v
16
/.0_v .20 u
a
c
-4 0.
0
//j -4
u
0Q
.2
v '^
L/D
o /280
8 0 /00
o
e
Airfoi..• N.A.CA. 4521
R.N3
: /50,000 /2
Dote: 4-17-31
Test: VD.T 573 _16
Corrected to infinite ospect rotib
V u
24 0 20
4 uu
_2
.4
/.6 .32
:/7
ROd _6T_
4g16D60
0
C p -40
v
-80
c
2.0 .40
0
5/ape cf "of
through end of
chord: 4/30
20 a 40
00
,44
1-T
0
0
qtu
.3
0 -.4
.2
QD /0 ITT
i 0
-/O
O 20 40 60 60 /C
Per cent ofchord
70
60
5.66
90 323
100 !./2) (-.%21
/o0 - 0
k
a
v
B
ou
FIGURE 48.-N.A.C.A. 45
L
°
0
Angle of otlock, a (degrees)
O
2
02
0/
-6 -4 0 4 8 /2 16 20 24 28 32
Up ^.
2.26
3./3
4.36
S.
6.02
7.17
6.0/
8.60
90l
936
9.16
6.56
p,
.Bo .lso
8
.6' ./2
w
4
.4'.08
0
J
0 0 y
0
.2 .04 VE
Airfoi/.• N.A.C.A.46/2 R.N.:3,2/0,000
Si­ 5"x30"
Vel.(fl./sec):68.3
P es.(slnd.olro):21,0 Dole: /2-/7-3/ -.2
Where tested.-LM
A.L. Test: V.D.T. 734 _4
Corrected for tunnel-wall effect
0 4 B l2 /6 20 24 2B 32
Angle of ottock, a (degrees)
v
16 p
0 .03
.4 .08
2 .04
Airfoi/: NA.C.A. 452/ R.N.:3,150,000
6/ie: S°x30"
Ve/.(ft./sec.J: 69.0
Pres.(sfnd.otm).•20.6 Dote: 4-/7-3/
Where tested:L.MA.L. Test: V.D.T.573
Corrected for tunnel-wall effect
.04
v 0
u
ev.160
o O
Co
S^.
/.25
2.5
SO
7,5
/0
15
20
25
30
40
50
zo
v $
L2 V .24 v
c
Gwr.
-1.57
-2./6
-2.81
-3.17
-3.40
-3.56
-3.50
-331
-3.00
-2.27
-/.4/
- .56
./0
.40
.3/
0
l
24 p
Ci u
.6^./2
-6
l
o,
1.6 .32 0 06
Ci
L/D
0Q
c
-4
40
.09
.44
2.0 .40
0/2280
60100
0 4u
44
V
E
co
P
./0
LO
0
48
.//
20 40 60 80 /C 7
Per cent .{chord
c. p.
16 -60
L
U /O
ui 0
29
Airfoi/:NA.C.A. 4612
Dote: 12-17-31
0
R:N:3,210,000
Test: V D. T. 734 -l(
Corrected to 1/7fhffe aspect rotio
-4
F-it. 49.-N . A.C.A 4612 airfoil.
Lift coeff/cie,L Cc
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
30
St . up L--
48
No /0
/.25 221 -/.6/
253.04 -222
50 4.24 -292
10 5.13 -3.32
U t 0
y k° -/0
44
O
l0 5.64 -3.56
/56.95-3. 79
20774 -377
256.3/ -3.60
30 9.71 -156
40 9.06 -2.56
50 6.
- /.64
60 6.47
47 - .65
20 40 60 80
Per ceniofchord
v
0
44
70].6] .34
BO 6.21
.95
90 3.73 .78
95 02) .42
100 L/2) (- 0
/00
0
L. E. Rod.: 1.56
Slope ofrodLs
0
U
/L
40
09
36
2.0 .40 Uo .08
N
18 .36 .0 .07
Ci
32
16 .32 0 .06
28
24
V U
endof
28 D 0 ihraugh4/3s
chord.
/4 .28c
24 0 20
A2,j. 24 b .04-/6
°.OS
v '^
20 0 40
l
p 16 ^ 60
L
0
o / t 80
/.Oy.20°
u
.03
8
.12
.0/
4
.4 .08
0
0
0
k
.0
0
-4
.2 .04 y`
n
v
0 0 y
Airfoil: NA.C.A. 4712 R.N:3,160.000
u
Si-S"x30"
1/e/(tf./sec): 68.7 2
.3
Pres. (sind.oim):210 Dote:12-18-31
c
Where fested'L.MA.L. Test: V.D.T.735 _.4
_
E
Correctedfor iunnel-wo//effect
' 4
C
0.
-8
-8
-4
0
4
B
/2
ao
^4
.02
.6
6 °loo
0 40U
4
20
v ^
3
/2
/6
20
24
Angle of ottoc/, or (degrees)
20
Z- toil. IV. A.0.A. 4712
R.N.: 3, 160, 000 /2
Dote: 12-18-31
rest: VD. T. 735 lB
Corrected to infhfte aspect ratio
-2
£
32 -8
0
.2
.4
.R
A
CO
12
14
1R
4R
Lift coe fficienf, C
FIGURE 50.-N.A.C.A. 4712 airfoil.
Sfa. up'r. L'0,
0
-
0
/.2S 370- .79
2.5
4.99 -- ,79
.94
50 6.92
7.5 6.42 - .46
/0 9.56- ,15
/s //.09
20 //.74
/O
y O cu
U l 0
yy M.p 10
0
.26
44
20 40 60 80 10
Per centofchoro
7
Co, ;
/00
-
2.0 .40 G0 .08
32,
.36
.07
28
1.6 .32
0 .06
24'
C
80 5.30
.02
90 2.99 - ,03
03
95 1.55 -: loo (,/2) (-.",
O
(L
/.8
O
L E. Rad.: 15B
5lope ofrodi3s
v
V
end of
28 p 0 ihro39h
chard: 6110
24 kk 20
^o
20 40
V
o /6. 60
12,80
1.4 .28 0°.O5
04
c
-4 0.
-8
20
N ^
L2V.24- v 04
c.
/6
v
LID.
40i .20 o .03
/2
h
u
.8 tE ./6
a
.02
8
.0/
4
o
.6u
8 0 /00
° 4u
.09--36
.44
50 /0.49 :.12
60 9.//
.05
70 7.37 .0/
a
40
0
25 //.92
3011
.03
- . /O
4011.
-./7
°C
48
./2
.4".
n
Airfoil: N.A.C.A. 6212 R.N.:,^240,000
Vel.(ft./sec.): 66. l
Size: S"x30"
Pres.(st'nd. ofm):20.9 Dote:12-21-31
Where tested: L.M.A. L. Test: V.O.T. 737
Corrected for funnel-wall effect
0
0
-41
0 0 y _.2
u
_4
-8 -4 0 4 8 12 /6 20 24 28 32
Angle of offock, a (degrees)
06
.2 .04 Ci
.3
0 -4-
c` -4 -2
FIGURE 51 .-N.A.C.A. 6212 airfoil.
_6
Airfoil: N.A.CA. 6212
R.N. 3,240.000 -/2
Dote: /2-21-3/Test: V D.T 737 _16
Corrected to infinite ospecf ratio
0 .2 .4 .6 .8 LO l.2 /.4 1.6
Lift coefficient, C
/.8
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL sf.. UP r L r.
125 1.63 - .40
2.5
v ^ /O
l
4.67
SB0
7.26
8.23
8.80
9.00
6.76
6.17
720
5.90
80 4.27
u
k
40
60
80 100
Per centofchoro
2.0 .40 x'-.08
48 .36 ^.07
0
L.E.Rod.: 0.40
5/ape of'oo-
G U
c
1.4 .2B
C.P.
u
40y.20, k'.03
u
o
k
a
.6.12
/2U 80
w
B 0/00
.o
12'
O,
.8v.16o .02
0 /6 60
4
20 w
/6 0
0.05
.v
/2V.24-Z^ v.04
v `^
q 20040
0
32
28 tt
1.6 .32 0'.06--
end of
0 lhra3qh
chord: 6//5
LID
y
36
8
.43
(-.06)
24 ^k 20
./O
.09
.44
.86
1.24
l00
v
0
20
0
100 (06)
- O
0
28
.50
.96
/.79
2.45
2.64
3.00
2.98
2.86
2.62
2.21
1,63
90 2.34
95
y 0 -/O
.OS
75
to
/5
20
25
30
40
50
60
70
0
u
246 -.34
50 3.79
31
.0/
k
4^
0
0
0p
-.2
-4 0
ro
-8
.4 J .08
C
W
g
.2 .04
0 O
u
v
Airfoil: N.A.C.A. 6306 R.N:,$080,000
Size: S"x30"
Ile%(fr/secf:69.8
P es. (st'nd. otm.):20.6 Dofe: 4-/7-3/ -.2
c -.3
Where tested.• L.M.A.L. Tesf: VD.T. 575 -.4-.4
Corrected for tunne%wol/effect
02
C
-4 0.
-8
-8 -4
0 4 8 /2 16 20 24 28 32
Ang/e of aHo'k. a (degrees)
-4 -,
AirfOi/:NA.CA. 6306
Dote:4-17-31
R.N.:3,080,000
Test: V D.T. 575
Corrected to infinite aspect ratio
O .2 .4 .6 .8 /.0 /,2 44 /.6 /.
Lift coefficient,C
FIGURE 62,-N.A.C.A. 6306 airfoil.
20
Slo.
S
U
0
U
0204060BO/O
0
Per Gent ofchord
.44
N
28 0 0
3
24 0 20
q20 0 40
chord.'
C
40.
U
-8
360
h
l
32
W
L8 .36 •u.07
C
28 tl
/.6 .320'0.06
u
1.4 .28-, p.05
15
v$
C.P.
L/D
/. 2V.24 ,4y v.04
160
v
c
v '^
/Ov.20a x.03
U
^^
O
^
.02
o /2 ^ 80
o
80)00
u
^ OQv
40
2.0 .40f
p 16 D 60
l
l
4
./0
.09
8
/00
a
00
2
/O
O
/.25
-.75
25
5.0 r
0S 76
7.5
50
/0
./B
/5
.46
20 9.70 1.02
25 10.29 1.35
30 10.50 1.50
40 10.24 1.53
50 9.50 1.55
60 6.35 /.4B
70 633 /29
60 4.94 . 50
90
.50
90 2.71
95 /.45
23
cos) !-0
loo o
L.E. Rod.: 0.89
Shope ofrad'us
/hrough end of
.2
0
Airfoil: NA.CA.6309 R.h!:3.110,000
Vel.(R./sec.): 69.4 -.2
Size: 5"x30"
Pres.(sihd.olmJ:20.B Dole:4-18-31
Where fested: L.MA.L. Test: V.D.T.576 -.4
Correc led for funnel-wo//effect
-8 -4 0 4 B 12 16 20 24 28 32
Angle of otlock a (degrees)
C
.0/
0
.04 V -./
0 y -.2
O
V
^-.3
0 -.4
`C -.4
6C
O
4^
.6i .12
.4^ .08
FIUVSE 63.-N.A.C.A. 6308 airfoil.
u
0:,L,
0
-4 0
v
4-18-31
Test: V. T. 576
Gted to fnfnile aspect ratio
.4
".6
.8
/.0
Lift coeffi'cient,C
1.2
/.4
16
1
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
32
St... Up.'r L'w'r.
-6---
1b /0
/.25 3.05
25 4.20 -1.38
5.0 5.93 -1.56
Ll 7.26 - /.47
yk y0
40
0
/0 9.36 -/. 29
15 /0.03 - .83
20
//. 14 -40
25 /1.79 14
-.
30 /2.00
/00
44
.08
60 9.50 .35
70 7.76 .39
90 5.62 .34
90 3.08
95
/00
1.66
l00 (./2)
o0
a
W
28 0 0
3
24^ 20
do
40
09
36
.08
32,
W
./S
.03
(-./?j
44
G$
2.0 .40
C
G. C. Rd.: 158
S/ape ofrod'us
yn endof
chord: 6/I5
C.p.
L/D
q
20 40 60 80
Per cenl of chord
.23
50 10.84
o x est No. 738
oo q
577
.//
.00
40 11.69
D
O
48
./2
0
U
1.8 .36 w .07
28
1.6 .32 0U
06
24'
1.4 .2,9-, 0 .05
20'
/.2V.24' w .04
/6
v '^
40 .208 o .03
200 40
0
o /sa 60
e`* ./s o
o /2 u 8 0
.6X ./2
w
.4J.08
/2:
8'
.02
0 0
8'-/00
c
0 4v
U
Ile
C
O
.O/
4'
0
0
-4'
.2 .04 G
u
ll
C. -8
0 0 w _2
i
Airfoil: N.A.C.A. 6312 R.N.:3.f80,000
ou
Ve1.(R../sec.): 68.5 .2
size: 5"x30"
Pres.(sfnd.otm):2/.0 Dote:12-22-31
c ' 3Airfoil: N.A.CA. 6312
R.N.:3,180,000
-/2
v _
8 u
Where lested: L.MA.L. Test: V.D.T. 738
Test: V. 0. 738 -/B
Dole: /2-22 3/
0 ' 4Corrected to infinile ospect rolio
Corrected for tunnel-wa//effect -.4
-2
n
.,9
LO
L2 14 L6 1.8
-4
.2
.4
.6
12
/6
20
F
24
28
32
-6 -4 O 4 8
Lift coefficient.0
Angle of attack, o: (degrees)
v
04
-4 0.
FIOVAB 64.-N.A.C.A.
Sic. up'r. LM,
/25 399 /.844
U
4 t0
20 /2.60 -1.63
25 /3.25 1.62
30 /3.50 -/.50
40 13.15 -/.37
5012.1660 r0.67 - .77
]O 8.7/ -.50
BO 6.30 - .30
0
/0
_/0
0
9.6] -236
i/0
/5 //.45 -2.13
20 40 60 BO
Per cen ,fcha d
v
28 0
-
0
C
L.E. Rod.: 248
5/ape ofrod'us
0 through endof
chord.' 6//5
c. p.
k
L/D
u
0
4
C
24
1.2x.24
/6
28
20
k
U
6
6^./6 0
O
O
.6X./2
k
o l2ru 8 0
8 0 /00
o
/.8 .36
/.6 .32 0
V u
1.4 .28 o
v
l
0 4v
o°
c
32
1.0 ,v .20
p /6^6D
l
36
U§
v
3
24 0 2 0
020;4 o
40
2.0 .40
/00 (./6J (- 0
/00
10
.44
90 a45 - .20
95 /.BJ-./J
'w
W.H.
0
25 5.15 -
SO ].OS -2.27
75 948 -2.39
"3
U
e:
0
.4
C
a°
4
0
.08
.2 .04 G
4
0 0 m
-8
Airfoil. N.A. C.A. 6315 R.N:3,100,000
U
Size: 5"x30"
3el(0.1sec.): 69.7
-4 4
Pres. (surd almJ:20.5 Dote:4-20-31 2
c
_
L.M.A.L.
Test:
V
DT
578
Where
tesled.
_4
Fc
_
_8 v
Corrected for tunnel-wo/l effect
-4
-8 -4 0 4 8 12 16 20 24 28 32 Angle of allocic, w (degrees)
FIGURE 56.-N.A.C.A. 8316.11U1.
/2
Airfoil. NA.CA . 6315
R. N: 3100,000
Test: V D.T. 578 -16
Dole: 4-20-3/
Corrected to inf nice ospect ratio
2
4
R
R
/0
/2
Lift coefficient. C
14
/R
IR
33
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL w o /0
u^O
o -/0
0
2
20 4060 60 /00
Per cent ofchord 1
./0
.44
'z.e2
O
0
V
1.40
2.0
5
45
0
LB .36
.5a
Y'us
16
C
5f
28 0
i
Bu
D
ti
aC
I
Ito. Up'r. L'Owr.
/.25 5.70 -1.79
2.5 7.20 -2.65
9.36 -363
5.0 11.03
7.5
-4./5
l0 12.35 -444
l5 14.26 -454
20 15.54 -4.65
-4.58
25 /624
30 16.50 -450
40 16.06 -4.25
50 14.64 -3.7/
60 12.99 -3.02
70 /060 -230
80 7.69 -1.60
90 4.22 - .90
95 2.31 -.55
( 22' ( 02)
%00
L. E. Rod.: 4.85
5/ape ofrad LS
through end of
u
C
0 Qv
C
-4 0.
-8
rest: V. D. T. 579
Corrected to k7finite aspect ratio
l
.2 .4 .6 .8 l.0 1.2 1.4 1.6 1.8
Lift coe ffcienf,C
.44
.05
h
36%
2.0 .40
G$
•.06
32 D
1.6 .32
C
l
28 21
d
24
's
0.06
l
20-.v
/6 p
v
1.4 .28-, 0.05
va
420.240 m.04
cp
10y.20
m `^
.03
2c
.0
8010
4
_/2e
Dole: 4-20-31
v
LID
^
v
_
1.8 .36 . 0.07
o /2 B
: o
0
-4 0
.//
i
o 16D6
U
03
20 40 60 80 /0
Per cent ofchord
chord: 6//5
24 wa 2
q 20x4
y ^
1, o /0
ul US O
0 -/0
0
4
4y
Y
2 .04
k
-40.
k
.0/
0
0
O U 2
Airfoi/: NAC.A.6318 R.N:3080,000
Size:5"z30"
V.Z(ft/sec.): 69.3 _
.3
P es. (sfnd.otm):2/.0 Dofe:4-20-3/ ' 2
Where tested: LMA.L. Test: VD.7.579
.4
.4
Corrected for funnel-//effect --4
48 12 16 20 24 28 32
-.4
_?91' of attack. a (degrees'
Fcccaa 56.-N.A.C.A. 6318 airfoll.
C
160
v
l2c
4
U
.8^.16^ .02
0
0
n
DC
0
0,05
6Xk .12
.4' .08
U
W
C
0
v `^
l.Ov,208 11.03--
L/D
4N
28 0
a
0.06
U
my
90
0
.32
L2V .24'M x.04
c.R
0 /2
.0
U
D
v
1.4 .28
3
24 a
k
q 20 0
0l 16
v
..08
e
0
D
v
.40
.36
-./9l
h
.09
C
Airfoil.•IV.A.CA.6321 R./V.:3/40.000
Si- 5'x30'Uel(ff./sec.): 69./
Pres. (stird.otm):20B Dof-4-21-31
Wherefesd:LMA.L.
rest: V. D.r5B0
fe
Corrected for tunnel-wol/effect
8c
.BV./6o .02
0 0
0
.0/
4^
.6,Ù .12
0
.4' .08
0
o^
0
.2 .04
-4 0
w
0 0 -.2
-8",
0u
Q
-.2
^ -.3
R.N:3,140,000
AirfoiLNA.CA.6321
v
7ez f: V. D.T 580 1_
4-21-31
-.4
0-.4 FTT- R Dote:
Corrected to infinite aspect ratio
-8 -4 0 4 6 12 16 20 24 2B 32 Angle of atfacl, a (degrees)
-. 4 -2
FrooaE 57.-N.A.C.A. 6321 airfoil.
0 .2 .4 .6 .8 LO 1.2 /.4 l.6 1.8
Lift coefficient C
34
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
Se. u,.-,. L'W^.
Z25 /.45 -.52
2.5 2/6 -.55
SO 3.32 -.36
7.5 4.24
zD /0
0
L,4
y p-/0
-.06
20 40 60 80 /6
w
W
28 p 0
3
041, 20
w
1.61
44
.07
28
1.6 .32 0U
06
24'
/.4 .28- 0
.0$
20
•v a
1.2V .24
.04
/6
LOV.20u^ .03
/2'
v '^
A
4^ 12 u 80
8 0 /00
./6 0
.6U./2
4^ .08
.2 .04 C
C
04
c
-4 0.
-8 u
Airfoil: N.A.C.A. 6406 R.N.:3,100,000
Vel. (ft./sec.J: 69.2
Size: 5"x30 •'
Pres. (stud. ofm.): 2/.0 'Dote: 8-31-31
Where tested.• L.hLA.L. Test V.O.T. 65B
Corrected for tunnel-wolf effect
S1a. up}. L'wY.
(25 206 -, 0BB
252.96-L//
5.0 4.30 -/./B
5.421 1/0 .66
-/
/5 7.78 -.36
20 966
25 965
3010.13
40 /035
24
4'
0
0:
./7
69
C C /O
U•C
O
O
20 40 60 80 /0
Per cenfofchord
%/2
665
.44
SO 9.B/ 486
60 8.78 1.92
70 7.28 L76
80 5.34
90 2.95
95 /.57
/.36
.74
35
/00
O
2.0 .40
C,
/00 /.091 (-.09)
d
/.4 .28 c
chord: sFzo
.w
20
420040
p /6
1.8 .36
1.6 .32
L.E. Rod.: 0.89
S/ape ofrodius
lhrough end pf
p.
/.2V .24
40 . .20 u
60
A./6
./6
£'2 71 80
.6 ° .12
8 0 /00
o 4m
.4-'.08
.o
O
O
.2 .04
°
l
C 04
c
-4
8
u
u
0Oo
i
u
-.2
Airfoi/:
N.A.CA.
6406
R.N:
$
/00,000
-/2
v _ 4Dote: 8-31-31
Test: U. D. T. 658 _16
-.4
Corrected to infinite aspect ratio
40_/0
3
0
8
O/
-8 -4 0 4 8 12 16 20 24 28
32
.4 -2
Angle of offock, a (degrees) F-.. 66.-N.A.C.A. 6406 W.H.
v
0
.02
S
kc
° 4 u°
.o
28 p 0
32.
c
/.8 .36
/6160
0
36
09
2.0 .40 U$ .0B
70 6.35 2.66
ao ass 2.oz
sa 2.59 /. u
95 /.3B 55
/00 (.061 (-.061
100 - o
L.E. Rod.: 0.40
Slope
lhrough
If-d",C
120
chord: 6/20
q 20 o40
^p
U
40
./0
Per centofchord
25 8.16 2./6
30 8.64 2.62
40 6.90 3.10
50 9.49 3.19
60 7.64 3.05
0
48
44
O
/0
506.2B
/5 &J9.97
20 7.42
.12
0
Airfoil: NA.CA. 6409 R.N.:3,060,000
Size: S"x30"
1/eL (f/./sec.): 698
Pres.lst 3 otm):20.8 Date: 9-/-3/ -.2
Where tested.• L.0 L. Test: l!D.T 659 -.4
Corrected for tunnel-wo//effect
-8 -4 O 4 8 /2 16 20 24 28 32
Angle of offock, a (degrees)
O
FILM. 69.-N.A.C.A. 64W .M.11.
0
.2
.4
.6
.8
LO
Lift coeffic/ent, 6i
L2
/.4
1.6
1.8
35
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 9Yo. up 'r.
/25 273
25 3.60
SO 5.36
73 657
to 7.58
1S
.A18
20 /034
25 11.14
30 11.65
Uw'r. C L /0
-123 U C 0
-1.64 h o_/0
-/.99
-,?.
0
20 40 60 80
-1.99
-1.67
Per cent ofchord
-1.25
-.76
- .38.20
4011.110
O
s0 /1.16
60 9.95
.65
BO a
.73
90 333
.39
9S /.79
/6
/./2) ('12)
/00
C
100
.55
.711
70 6.23
u
2
•//
./0
40
.44
.09
36 v
2.0 .40
0.08
4
32 D
0
C
L.E. Rpd.: 1.511
v
5 otrod/us
/ooe
end of
28
0 through
^hord:s/zo
LB .36 .07
28 t
.06
24 p
1.4 .28c O0 .05
v^
/.2.24
:0 v 04
ti
20,
1.6 .32 0
.0
l
v
16 p
Ci u
0
24 20
M
420E40
y o
LO 1.206
v
.03
12 c
U
e^./so .02
16 60
o
.6X./2
k
8 `0/00
k c
0 4m
.4 .08
-4
0
4
8
C ^,
-4 0
v
_ 8 0`
U
12
e .3
R.N.: 3,0917,000
Airfoi/NA.CA.
.•
6412
De:/2-22-31
Test: VD.T 739 _16
E0 -.4-1co _ -cfed to infinite os ect rofb
sf7dolm):20.8 Dote:12-22-31
Where tested•L.MA.L. Test: V D.T. 739
Corrected for tunnel-wall effect _.4
-8
0^
_, 2
O 0 v
-4 4
4
0
Airfoih N.A.C.A.64/2 R.N:3,090,000
VeG(f//sec.): 69.5 _, 2
Size: S
-B o
^
.0/
2 .04 C
C
'
0R
C
8
q
o /2g 80
.O
48
00. Test No. 739
000
660
-4 .2
/2 /6 20 24 28 32
Angle of attack, a (degrees)
0 .2 .4 .6 .8 /.O l.2 L4 L6 L8
Lift coefficienl,C
FIX 60. N.A.C.A. 6412airfoil.
$YO. Up'r. L'wr.
O
-
/.
225 345 -/.53
2.5 4.67 SO 6.44 -2.2.13
]5
8.776 -ao6
z9
/0
15 /0.556
o -2.97
20 /L6/ -267
25 12.64 -2.29
30 13.15 - L91
50 1246 - .76
BO //. /O - .34
706.70
9./6 -.04
80
.09
90 372 .04
95 2.01 -.05
D
20
./2
w p 10
-1
4,4
p
0
0
0
s
20 40 60 BO /0
Per cent of chord 0
0
44
40 1325 -1.25
r
ti
2.0 .40
V$
-
.09
W
v
.08
a
U
N
1.8 .36 .u.07
too 1'161 f-.16/
C
28
L. E. Rod.: 2.48
0
Slope ofra0 us
through end of
0 chord: 6/20
0
III
1.6 .32 0.06
Co
W
3
24 0 20
20 0 40
y
o k
/.0 .0 .20 0.03
L/D
l
/2U 6 0
0Q
c
-8
8'11
o O
.6,', .12
.0/
4s
0
Oo
.4'.08
o 4v
-4 0.
ti
.8'.16 0 .02
k
h 8 0/00
` C
u
/2'c
U
^
o
/6 p
v
1.2V.24^E- x.04
C. p.
16; 60
l
u
1.4 .28 0.05
0
-4 o
v
'B 0,
e
.2 .04 tj
a
Airfoil: NA.CA. 6415 R.N:3,060,000
Si-5"x30"
Vel(ft/sec.): 69.8
Pres.(sthdotm):208 Dofe:9-3-3/
Where fested.- L.MA.L. Test: V.D.T. 661
Corrected for tunnel-wolf effect
0
0 Q̀, -. 2
U
-/2
-.4
-8 -4 0 4 8 12 16 20 24 28 32
Angle of attack, a (degrees)
k
v
0 -.4
-4
Frovas 61. -N.A.C.A. 6416 airfoil.
Dote:9-3-3/
Test:
Corrected to infinite aspect
.2 .4 .6 .8 LO 12
Lift coefficient,C
36
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
0
0
ur / 0
v.; O
1. 25 4.26 -1.92
25 7.62 -2.59
3.0 7.53 -346
Cp
].5 9.99 -39/
l0 10.06
0
-4./5
05 1202 -4.26
20 /332 -4.0]
25 14./7 -3.75
30 14.64 -340
40 14.70 -2.70
50 13.60 -205
60 1224 -447
70 7.40
/0.1/--.94
60
.54
90 4.12 -
u
k
20
40
60
60
Per cent oFchord
/0
/0
h
.44
36 m
l
32 U
U^
2.0 .40
v
46 .36 .0
95 2.24 - .23
28 b
100 (./91 (-./9)
/00
b
v
.0
0
L E. Rod.. 3.56
Slope ofrodius
end of
28 0 0 1hro3qh
chord: 6/20
3
24 0 20
C
/.6
.32
U° U
a
.v a
16,
v
12:c
v k
44 .20 8
q20040
o 16 O 60
.0
p a
`c
Bv.16^
L/D
^°
0 a
.6X ./2
o /2^8 0
k
8 v l0 0
o
4r
0
00
.4 .68
a
n
0 4u
v
^0a
- 4 0.
8u
-8 -4 0 48 12 16 20 24
Angle of attach, a (degrees)
0
0
-
C
v0
0
2
c
_4
e
-.4
28 32
FIGURE 62 .-N.A.C.A.
Sta. up ". L' w 'r.
-4 0
2 .04 C
Airfoil: NA.C.A.6418 R.N.:3.100,000
Ve%(R../sec.): 69.4
Size: 5"x30"
Pres.(st'nd.ofm.): 20.6 Oate: 9-4-3/
Where tesled: L M.A.L. Test: V.O.T 662
Corrected for tunnel-wo/t effect
6415.W.11.
`
48
Ur
0
m 0 _/O
44
.//
0
/O 1/.52 -5.18
/5 /344 -552
4
20 / .79 -5.49
25 1565 -523
30 16.15 -4.9/
40 /6.16 -416
50 /5.14 -3.40
60 13.44 -2.59
70 11.06 -/.93
80 8.06 -1.17
90 4.51 - .65
'95 246 - .42
20 40 60 80
Per cent ofchord
t0
40
36 y
.09
44
2.0 .40 ­ 08-
32 D
v
28 8
18 .36 .0 .07
.d
24 0
loo (.221 (-.221
/00
O
L.E. Rad.: 4.85
U
W
AS .32 E3,06
u
Slope
of
ughradiuI
end of
28 p 0 Thro
chord: 6/20
24 0 20
w
y20 0 4 0
616-1-6
0
l ,
0
C
l
20 w
/6 0
v
1.4 .28 1 0.05
W
a .04
V ^
v k
LOy.20
03
.8x./60 .02
01
.O/
.6w .12
0
.4'.08
c. p.
L/D
a 12 ^ 8
B o t0tl
12:c
8.c
a
4
4^
0
00
Y
0 4.C
.o
.2 .04 G -. /
O 0 k0 -.2
Airfoil: NA.C.A. 642/ R.N.:3,030,000
u
Ve1.(fl./sec.): 70.5 .2
14EE: 5"x30"
.3
Prss.(sf'7d.olm.):20.4 Ooie:9-4-31
-4 0
C
0 Qv
.0
4
Where tested: L.M.A.L.. Test: V.OT. 663
Corrected far lunnel-wall effect -.4
U
-8
-8
-4
O
w
_8 0
Q
2
7: NA.C.A. 6418
R.N: 3.100,000
9-4-31
Test. VO..T 662 _16
tied to infinite aspect rcho
.4 .6 .8 10 12 1.4 L6 7.8
Lift Coefficienf.0
y o /0
O
/.25 5.13 -208
25 6.60 -3.04
5.0 5.65 -4.16
7.5 10.24 -4.81
S
k
20 u
1.4 .28-, o
42V.24;^ v
c.p.
U
24 p
l
0
48
12
/6
20
24
28
Angle of offpch, a (degrees)
32
Fa._ 4
0
_8 P
Q
Airfoit: N.A.CA. 6421
Dole: .9-4 -31
Corrected to infnits os
-.4 -2 0 .2 .4 .6 .8 10
FIGURE 63. -N.A.C.A. 6421 airfoil.
Lift coefficient. C
V O. T. 663
37
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL s .
1.25
2.5
50
7.5
/0
15
20
25
30
40
50
60
70
80
90
v
Up ^.
1.36
2.00
301
3.65
4.58
5.76
6.74
7.49
8.06
6.66
8.65
8.06
6.69
5.17
2.90
G' Y'.
0
- .60
- .70
-.62
- .43
- . l9
.37
.95
/.51
2.03
264
3.35
3.48
321
D /O
C0
0 _l0
O
./2
20 40 60 60 / 00
Per centofoh-d
L.E. Rod.: 0.40
,,, ,he
ndof
v
1,111
04 0 200
k
0120040
01660
36
.09
2.0 .40 Ci.09
W
/.8 .36.07
/.6 .32 0.06
L52
1.43
95 /.57
73
100 (.06) (-.06)
/00 - 0
U
./O
44
32
?6 ^
?4 p
l
?0
V u
chord: 6125
44 28C 0.05
C
LD
42, .24^u \y',.04
v
c. p.
1.q.2004.03
U
.d .160 .02
12 80
.6U.12 .0/
60/0
.4'4 .08
0
s
C
04u
.2
.04
yr
u
.v
0Q
0 0 -.2
0u
Airfoi/:N.A.C.A.6506 R.K:,$17Q000
C
Ve/. (R./sec.): 68.5 -2
Size: 5"x30"
-4 0.
e-3
Pres.(stnol-):2/./
d.
Date:4-23-31
Where tested: L.MA.L. Test. • V.O.T. 586 4
-8 v
Corrected for tunnel-wall effect
-6 -4 0 4 8 /2 16 20 24 28 32 4
Angle of attack, a (degrees)
/6p
4,
m
lC ^
8
-foil: N.A.C.A. 6506
R.N: $170,000 -/2
te: 4-23-31
Test: V D.T. 586 _/8
rrected to infinite aspect ratio
.2 .4 .6 .8
/.0 1.2 /.4 L6 1.8
Lift coefficient,'C
FIOME 64.-N.A.C.A. 8608 airfoil.
s11.
0
1.25
2.5
c
up'r. L'w'r.
- 0
/.93 - .99
2.75 - 1 .27
3.59 -445
v o /0
t 0
Ca0
50
7.5 4.97 -/.34
45
/ O 5.62 -L
O
U
0
a
v
O
/5 7.18 -.96
20 8.22 - . 9
3.56 .5/
30
4010.11 /.39
50 9.97 2.03
60 9.20 2.34
70 7.81 2.30
80 5.85 1.87
90 3.29 /.07
95 /.78 53
/00 (.09) (-.09)
/ 0
0
LT Rod.: 0.89
5/ape ofrodius
/hrough endoi
28 'b 0 chord: 6
3
24 0 20
20 a 40
V
p /8 1t 60
a /2 u 80
8"0 100
20 40 60 60 /G
Per cenl ofchord
.44
W
C
.36
u
16
.32
0
G °
25
p.
-8
-8
-4
8
4
0
0 0
.6u./2
.4'.08
Angle of oNock, a (degrees) O
4
8
x
.2 .04 C
0 0 y
Size: 5"x30"
Ve1.0../sec.): 69.4
Pres.(51'ld. 11-120.6 Dote: 422-3/ -.2
Where tested.• I'll"L. Tes1:VD.T.585
Corrected for funnel-wo/l eff ct -.4
-4,
/6
/2
u U^
6 /6°
Airfoil:NA.C.A.6509 R.N.:3,110.000
.c
l6
1.4 .28-, 0
N U
L2V.24 ^y v`
/.Ov.20^ o
C
°o 4uU
OQv
V
2.0 .40
12
16
20 24
28 32
0
v
a
Dote:
-.4 -
Fmvaa 86 .-N.A.C.A. 8600 airfoil.
L 1ff coetfiaenf, c^
38
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
20
5fa LPL-.
a -
o
tzs 2n -/..la
25 356 -/.82
SO 5.02 -226
75 613 -2.43
15 857 -2.27
-1-91
25 /650 -1.47
4011.6- .06
so 229 7/
60 1035 i.21
70 876 L39
BON
6.5 I.24
RLI
1111
46
.//
w -/o
yo
Test M
20 40 60 80 /6
Per cenf ofchord
O
/O 7.06 -2.45
o a x
eoo
44
40
1
20 9.69
3011.07-
U
.44
.98
O
Ioo
c
W
28t D
24 ,420
^20i4D
o I6g 60
o
B0
C
-4 0.
-8
32
C
1.8 .36 .u.07
w
o
L. E. ROd.: I.SB
5roae ofrodis
28 e
.O
24 p
l
20 u0
/6 0
1.6 .32 0.06
V U
1.4 .2B 0.05
Through end of
chord: 6/25
•v D
.2 24 x.04
c.p.
L/D
v
vk
110
4.03
/2.c
v
a
.8v.1.6 .02
0 0
.68w ./2 .0/
.4'.08
Bolo0
°0 4u
0Q
36%
2.0 .40 G.08
90.72
-°C
d
.09
a
.2 .04
C,
u
0
0 w0 -.2
u -.3
v
-.4
0
Airfoil.• N.A.C.A.6512 R.N.:3,180.000
size: 5'x30"
Vel.(ft./sec.): 68.6
P es.(sfndolm):20.9 Dote: /2 23-3/ -.2
Where lesled: L.M.A.L. Test: V D.T.740
Correcied for lunnel-11 effect
-8 -4 0 4 8 /2 16 20 24 28 32
4ng/e of attack, a (degrees)
k
4y
F -4
0x
0
-4 0
v
-B
/2-23-31
sled to infinite as
.4 .6 .8 /.0
Lift coeff/cieni,C
Fiaass 66 .-N.A.C.A. 8613 ailfb 1.
S^ OPr LW r.
125 3.26 -1.68
2.5 4.40-2.33
b
0
ZU
0
aC
N
D
28
S.0 604 -3.04
Z5
15 8.34-355
297 -3.56
20 11.14
-3,33
2S12.00 -295
30 /259 -249
40 13.00 -1.52
5012. 2- .62
60/450
.0e
70 9.69 .49
80 7.22 .60
90 405 .36
/00 (:/5i (-JS)
/00 -
o
L.C. Rod.: 246
0 t ough ^dof
l0
U 0
4,4 _/0
p
0
20 40 60 80 /L
Per cent ofchord
.44
2.0 .40
/.8 .36
Ci
/.6 .32
CS
/.4 .281
chord: 6/25
v
3
24 L.
L2 Ci.24cvov
/.0 y.20°
c. p.
w
N20040
LID,
v
16 60
t
^
0
o /2 80
8'.100
a
° 4u
.o
n
OQ
C
.2 .04
0
Airfoi/..NA.CA.6515 R.N:3,100,000
Size: S"x30"
Ve1.(ff./sec.): 69.4
Pres.(sthd.otm.J:20.11 Dofe: 4-22-3/ -.2
Where lesled•L. .40 Test: V.DT. 583 -.4
Correcfed for tunneFwof/effect
v
-B
-6
0
.6 °./2
k
.4' .08
-4
O
4
8
12
16
20
24
Angle of attack, a (degrees)
0
28 32
FIOU88 87.-N.A.C.A. 8616 slRoll.
Lift coefficienf,C
Q
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL w^ /o
u °c 0
39
./2
48
EFFF
C o -/0
44
0 20 40 60 60
er cent ofchord
p
40
zss
Ls4
Los
.41
a
05
U
S
0B1
56
dlls
Sf
v
m
28 0
24 0
44
^0
/,
2.0 .40
1 .08
w
/.B .36
C
N'0o
^
24
ZO
x.04
k
^.03
/2
02
C
-8 u
5^ . UP'n L'^ r.
/.25 4. ]5 -222
2S 6.17 -3.26
SO 6.20 -450
9. ]0 -524
75
/o 10.93 -57/
/512.79
2014.10 -6./3
-.6.14
3015.60
25 /505 -59Z
-553
-444
4015.63
5015. 27 -3.27
601181 -220
1457 -/.30
70
80 6.59 - .6B
90 4.B5 - .32
95 2.67 - .23
(22) (:22)
%000
0
L.E. Rod.: 4.65
6/ope of radius
1 / eo endof
chord: 6/25
u,
0
4
a
.2
.04
0
0
/6
0
.02
8'
0/
4
0
0:
v-.2
Cw-8
w_' 3
Airfoi/.•N.A.CA. 65/6
R.N:3,/00,000 -/2
Dote:4-21-31
Test., V.O. T. 582 -16
Corrected to infinite aspect ratio
o
'4
-4
..2
0
.2
.4
.6
.8
/0
Oft coeff/clent,C
/.2
/.4
/.6
t.8
68.-N.A.C.A. 6618 airfoiL
^ a ^O•
^1Q 0
R 0-/0
O Per
20 cent
40 ofchord
60 80
/6
/0
C
N
l
.09
36 m
2.0 .40 V^..06
32 a
/.8 .36 x.07
28
/.6 32 0.06
24 p
44
U° u
.o
/.4 .28 0.05
l
ED'.
/.2 V .24 v.04
/68-
v
24 l 21
c.p.
L0^.20U x.03
020041
/6^ &
o 12 1u Bi
B 0 /0,
0
4u
.o
04
_4'0.r
u
Airfoi/: NA.C.A. 65/6 1?.1:3, IOQ000
Size: 5"x30"
Ve%(ft./sec.//: 693 _ 2
Pres.(sti]dolm):20.8 D&.%1_31
Where tested.-L.M.A.L. Test:VD.T. 582
4
Corrected for tunnel-wall effect
4 8 12 16 20 24 28 32
ng/e of a//ock, a (degrees) r'.
z6
m
.4'.08
-4 R
28
0.05
.6'./2
u
k 8 0/
C
0 4 vu
o
l
32
0
°
0.06
/.0 4 .20
Pf
°p
-
36
14 .281
1.2V .24
II I
C
l.6 .32
0 16^
O
U
k
07
w
I
I
c. P.
k
t
/ 0, 1
u
R
.02
L/D
0
u
.6 U .12
.01
4^
4`1 .08
0
00
0
40
.2 .04
v
0 Q 0 -.2
Airfoil: N.A.CA. 6521 R.N.:3,080,000
Size: 5"x30"
Vel.(fi!/sec.): 69.8 _.2
Pres.(sfnd.ofm):20.6 Dote: 4-21-31
Where tested: LM.A.L. Test- VD.T 58/ 4
Corrected for tunnel-wall effect
-8 u
-8
-4
0
4
8
12
16
20
24
Angle of oiiock, a (degrees)
28
32
U
C_3
-4 -
Airfoi/. N.A.C.A. 6521
Lift coefficient.
Fmm. 68.-N.A.C.A. 6521 M doiL
R.IV:3,080,000
Dote: 4 2/ 3/Test: VAT 581
Corrected to infnite aspect ratio
0 .2 .4 .6. 8 /.0 12 1.4 16 l
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
40
S" UPr. L'11
0
/.25 2.45 -1.43
2.5 3.41 -L94
So 4.78 -247
75 5.64 -270
/0 6.70 -278
15 &// -269
20 9.17 -2.41
25 9.95 -200
3010.51 -1.5/
4011.14 -.49
50//./3
so/
53
0.56 1.44
70 9.33 L98
0
BO 7.19 /.90
U
90 4./6 /./9
95 2.28 .60
/00 (./L (-.l2/
/00
O
c
L.
C. ROd.: L58
N
Slope Ifrpdius
/h
end
.28 p 0 chord: 6130 of
3
24 L.
Q
LID
^/
M0
to
Co
O
20
40
60
80
/0
Per cent ofchord
.44
2.0 .40
C
/.8 .36
1.6 .32
1.9 .28 C
,d
L2V.24
p.
20 0 40
/6 60
o /2.g 80
q,-/00
c
0 4v
o
/.O y.20
rJ
o
.6'./2
.4'.06
ix I
C
.2 .04
u
N
R 04
C
-4 0.
Ci
-8
v
0
0
Airfoil: N.A.C.A. 66/2 R.N.:.1250,000
Vel.(N./sec.): 68.0
Size: 5"x30"
s 7d.otm.): 21.0 Dote:/2-28-3/ -.2
Pres. W
M Test.• D.T. 741 -.4
Where tested.• L.A.L.
Corrected for funnel-wall effect
-8 -4 0 4 8 l2 16 20 24. 28 32
Angle of ot/ock, a (degrees)
Fl°II88
Lift coefficient.
70.-N.A.C.A. 6812.11M.
44
U
40
0
s
u
0
36
ti
U
C
v
28 p
3
24 0
k
.32
V
'28-,
v
V .24•^
w
v
v.20a
N 20
0 16 ^
u p,
v./60
O
O O
o /2 u
k
^
u /2
V
.08
8°
.04
° 46
v
OQ
C
-4 0.
0
U
-8
Litt coeff/'aent,C
Angle of oltock a (degrees)
F1Qv v 71.-N.A.C.A. 6712 e4(oll.
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL
41
cD
u
-2.55
Per cent ofchord
2.62
-296
-3.00
.44
287
a
-ass
-2.1s
-1.70
s
u
- /.22
0
k0
D
cv
28 ^
3
2.0
40
/.8
36
_ '33
(-.061
o.o/o
1.6 .32
ti
/.4 .28`C
v
/.2 V .24
c.p.
24 0
k
q 20p
L/D
LO• w .20 u
C
'u
6 0,
w
o° 4 u
^
0
0
04
Airfoil: N.A.C.A.0006T R.N.:3,180,000
uel.(ft./sec.): 68.7 -2
Size: 5"k30"
Pres.(sti7d.o/m):20.8 Doie:4-8-31
Where fested•L.MA.L. Test:VD.r 554 _ 4
Corrected for tunnel-wall effect
c
-4 0.
8v
0
4
8
12 /6
20 24 28
Angle of attack. a (degrees)
w,
/.42
1.65
2.30
2.54
2.69
2.87
296
3.00
3.00
296
2.79
Lm,
- q2
-1.85
-2.30
-2.54
-269
-28]
-2.96
-3.00
-3.00
-2.96
-2.79
/0
/5
20
25
30
40
50
60 2.49-2.49
-204
i
L ifl coefficient, C,
90 .77
95 4/
ERR
O
./t
20 40 60 80 /0.
Per centofchord
.t0
.41
!8
.36
W
^.07
!6
.32
0.06
1.4 .26
28 ^ !
3
24 a 2i
no
.09
.44
'.0 .40 1 .08
tD0 (.06) (-06)
too 0 0
L. E. ROd.: /.19
cN
-N.A.O.A. 0006T airfoil.
./2
u c° 0
70 2.04
80 /.46 -/.46
-.7]
-
0
0
32
Ftoo z72.
6fo.
/.25
26
S.0
7.5
o•
Bv.16o
0
0
.6 0 ./2
G
4' 08
2 .04
0/60
a /2 u
.
v
y 2 i4t
Q
D
b
^q U
,v
0.05
!2V.241^- m.04
/6 8
LOv.20 0
/2
U
a
.evo .lspO .02
o
80/0i
k
c
a
° 4uL
.o
Oy
.c
.A. 00068 RN.:3.090,000
"
M(H./sec.): 69.5
tm.):20.8 Dote:4-8-3/
-4 R
-8 u
d.L.M.A.L. Test: l!D.T.556
or funnel-wall effect
P
- a -4
V 4
c4 ca Jc
Angie of oHccA. a (degrees)
W
M
.0/
.6 ./2
0
.4".08
x
2 .04 G-./
0 0 0 -.2
°u
'2c .3
'•4-4
.4
4,^
0
Cw
-4 0
v
-6 a
e
Air. foil. N.A.C.A. 0006B R.N.:.X6
Test: Y 6
Dote: 4 -8 31
Corrected to infinite aspect ro:
-2
FlomtE 76.-N.A.O.A. 00068 abfolL
0
.2
.4
.6
.8
l.0
Lift coefficient, Q
12
14
556
42
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
sfo. upr. cwr.
a "
1.25 -12s
lU
0 -/0
0
125
as
5.0
7.5
/O
/5
20
25
30
40
50
60
70
80
90
95
/oo
Lao
0
0
/o
L88 -188
266 -gas
3.6/ -361
4.21 -421
5.09 -509
S.
5.64 -.464
-5.92
6.00 -600
5.75 -5.75
5// -5.//
4.29-4.29
339 -339
2.43-,.43
f.37 -137
78 -.78
i
(a2
0
0
L.E.ROd: 0.40
u
0
v
4
0
44
20
40
60
/L 7
80
40
Per cent olchord
1
.09
360
2.0 .40 <.08
32 U
44
W
1.8
V U
0 40
• a
42U.24* -W.04
m '^
1.01.20$ x.03
u
u
.8 .16
.02
a
0
.6'.12
.O/
L/D
U 16 60
12$ BO
0
.o
4
V
0y
-8
c
4^
U
0--.
.4.08
x
.2
.04
C
k
Airfoih N.A.CA. 00127 R.N.:3.120,000 0 0 0U -.2
/00
.c
_4 4
ti
28 i
0
24
l
20 m
/6p
12^
w
/.4 .28 < 0.05
c.p.
24 ^ 20
8
.36
16 .32 0.06
28 0 0
V20
48
Si- 5"x30"
Ve/.(fL/sec.): 69.5
P es.(sti+d.otm):20.5 Dole.' 4-3-31
Where ies7ed L.M. A.L. Test: V D.T. 548
Corrected for tunnel-walleffect -.4
-8 -4 O 4 8 12 16 20 24 28 32
Angle of attach, a (degrees)
v
1-.4---.4 .2
O^
0
-4 0
v
e
Dole: 4-3-3/
Test: U. D. T. 5481
Corrected to inf'niie aspect ratio j.2 .4 .6 .8 /.O 1.2 44 L6 1.8
Lift coefffcient.0
Fioasa 74-N . A.C.A. 0012T SW0.
125 ze4
25 3.69
50 4.59
75 S.5.08
10 S.
f5
20 592
25 6.00
30 6.00
40 593
50. 5.59
60 4.07
4.98
70
80 2.91
90 1.55
95 63
100 ( ..12)
-1 e4
-3.69
-4.59
-5.06
c l
/o
y
Ra
-10
0
-536
-574
-5.92
-600
-6. 0
-5.93
-559
-4.96
- 4.07
-2.91
-1.55
-83
(-.. /2)
loo 0
0
L.E.Rod.: 360
0
u
0
D
v
2
0
44
20
40
60
80
Per cent ofch.rd
1C
./0
.44
.09
2.0 .40
V$
- .08
f.6 .32 0U t06
1.4 .28-, o .05
ma
/.2V .24* v .04
v ':
401 .20 0 .03-.8x./60 .02
0 0
L<
c.p.
20040
0 /6 D 60
w
.4^ .08
0
4
4 0.
-4
0
4
8
/2 16 20 24 26 32
Angle of attach, a (degrees)
28
24
20
6
/2
°
8'
.0/
4
0
0.
C .ro
-4
0 0 m -.2
Airfoil: N.A.C.A.00128 R.N.:3.060.000
oU
Ve/.(ft..Isec.)_ 70./ 2
Size: 5"x30"
Pres.(sthd.otmf:20.3 Date: 4-3/
c '3
Where tested: LM.A.L. Tesf: V. D.T.550 _4
-4
Corrected for tunnel-waf/effect
-8 v
36
32
.2 .04
Ca
.c
40
CPo
W
24 w0 20
-8
7
/.B .36 .0 .07
28 Do 0
a /2 u 80
w
8 /00
^. c
0 u
v
0V
48
-
-8'
A/rf0lL• NA.CA. 00/28 R.N.:3.080,000 -I2
Dole: 4-4-3/
Tesf:V. D. T. 550 _/6
Corrected to Infinite aspect rotio
-.4 :2 0 .2 .4 .6 .8 40 /.2 44 l.8 1.8
Rama 76.-N.A.C.A. 00128 airfoil.
Lifl coefficient,C
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL
Sro. Up'r
25
SO
Z5
to
/5
20
25
30
U
0
L'wr
U
283 -2.83
4.28 -4.28
5.4/ -.£4/
6.32 - 6.32
164 -7.64
8.46
B.88 -8.88
-9.00
9.00 -846
R p -/D
0
20 40
60
80
Per centofchord
40 662 -8.62
50 767 -761
60 6.44 -6.44
.44
2.0 .40
L. E. Rad.: O.B9
/.8 .36
/.6 32
1.4 .281
v
70 S09 -5.09
B0 3.65 - 365
90 205 -205
95 /./7 -1.17
l00 (./B) !-.7Bi
0
/o0 0
t
U
o
0
c
N
/O
C
z8 ° 1
c. p.
24021
k
2,j. 04
v
LID
V 20o4i
o /6 &
/.0120U
u
8 = ./6 v
0 0
.6X./2
.4 .08
t
^
a /2 t 8
r
a
U
8,/0
4v
0
O U
.2 .04
C
0
04
C
Airfoi/NA.CA: 00/BT R. N:3,/50,000
Ve%(ft./sec./: 69.2
Size: S"x30"
Pres.(sfndatin):20.6 Dofe:4-7-3/ -.2
Where tested:L.MA.L. Test:VOT 552 -.4
Corrected for tunnel-watt effect
4 0.
u
-8
0
-8 -4 0 4 8 12 /6 20 24 28 32
Lift coefficient,C
Angle of attack. a (degrees)
F/evas 78.-N.A.C.A. 081ST slrfolL
^a
i^
30 sod
40 8.89
50 6.36
60 Z47
70 6.11
80 4.37
90 2.32
95(.78)
1.24
/0o
0
.0
U
4.
o
/00 0
a
0
-9.
-6.69
-BAS
-7.47
-6./1
-437
-2.32
-1.24
(-.16)
,44
2.0
.40
/6
,36
0
L.E.Rad.:]. /5
/,6 .32 q U0
V
1.4 .28%
y
L2^.24
C
ZB O
24 l 20
ap.
411
L/D
N 20 0 40
0 /6a 60
12. 80
00 6 0 /00a 4
40,^ .20
P
.e Q)
O
o
.6U./2
.4-.06
C
.2 .04 C
.o
p4
-6
,N
U.
0
0
y
0
Airfoi/: N.A.CA.
00188 R.N.:3,180,000
Size: S'k30 1,Vel-(ft/sec): 692 - 20
Prey. (sfnd. otm):20.3 Dote: 4 - 6 - 31
Where tested.•L.M.A.L. Test: V.D.T 551
Corrected for tunnel-wall effect _ •4
'C
u
-8
-4
0
4
6
12 16
20
24 28 32
Angle of attack. a (degrees)
-
Frayse 77. N.A.O.A. 0018B elrfolL
43
44
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
20
Sto up'r.
L.-r,
_
,. 0
57
2.53./0 -2./7
E /o
2
iO
yy,_/0
o
5.0 4.29 -286
75 5.16 -a28
/O 5.84 -3.57
15 6.82 -3.88
O
20 747 -402
48
44
20
40
60
80 / 00
Per cent ofchord
./0
40
.09
36 w
N
25 7.82 -4.06
30 796 -402
40 776 -3.84
50 7.03 -355
60 5.94
-3.16
4.61
70
u
.44
20 .40 c.06
/.8 .36 ll.07
i
-2.72
BO 316 -2./0
90 1.63 -4,26
95 .87 - .74
/00 (.13) (-.13)
/00
0
L. E. Rod56
.: /.
w
Slope of ivdius
end of
28
0 through
u chard: O. /3/
0
rz
C
24, 20
L6
k
L/D
o /6^ 60
12280
6 -/00
°o 4u^
C
04
Airfoil: NA.CA.2R/2 R.N.:3,260000
Size: 5°x30"
3el.(f1./sec.): 68.3
Pres. (sthd. otm):20.6 Dote:2-26-32
Where tested.' L.M.A.L. Tess: VD.T. 764
Corrected for tunnel-wo//effect
.c
- 4 0.
u
-8
-8
-4
0
4
8
12
0
24 0
l.4 .28.OS
42V.24 i^o y.04
C
/. 0,§ .20 ^.03
.A.16
.l6
.02
.6.120.0l
20 m
/6p
124-c•
8'^
.4".08
_
/6 20 24
Angle of atlock, a (degrees)
28
l
32
28 21
o.06
C
c.p.
q 20Q40
.32
L"a,
$
4
LLJ
0
.2 .04
0 O w -.2
u
-.2
-•3
_.4
_•4
u
°
C,H
-4 0
v
-g a
e
Airfoil.' N.A.C.A. 2R, 12
AR. 3,260, 000 -l2
Dote: 226 32
Test: V D. T. 764 -l6
Corrected to infinite aspect ratio -4 -2 0 .2 .4
32
0
1
.6
.8
/.0
Lift coefficient, C
12
/.4
1.6
1.8
Fro- 78.-N.A.C.A. 211 1 I2 airfoil.
20
Sta. Up'r. L'w'r.
O 230
- -/.52
OvO Fk /0O
JS
2.5316 -2.10 4w-/O
-276
S''
' -3.17
7.5 5.29
0
10598 -a42
156.97 -3.74
20
Z56-3
25 791
-3.97
30800 -400
48
44
20
40
60
80
Per centofchord
00
./0
44
.09
40
36
506.73 -.867
D
44l
60 S49 -a66
2.0 .40 -.oe
32 a
704. -3.27
O
60- / -2.64
W
U
90 1.26 -1.63
48 .36 ^ .07
95 .66 -.95
OOp
28
(.0 (-.13)
a 10
100 O
1.6 .320".06
8
24
0
L.E.itod.:
/.SBC
v
/^ou endof
28 0 0 !h
/.4 .28
.05---20t
chord: O. 153
C.P.
24 0 20
1.2V .24w y.04/6no
a
mk
L/D
20 p 40
1.0 .w
1 .20
o °x.03
u
V
u o C
k
a°
p /6D 60
.8./60 .OZ
8^
0
80
o
.6 .12
.0/
q^
u
8'-/00
.4'1 .06
0
00
=
o
0 4v
C .i.
.2
.04
vr
-.
/
-4o
3
o
u
0
v
_g a
0Q
0 O "Q:) -.2
Airfoil: N.A.CA. 2R2/2 R.N.:3,130,000
oU
Q
c
Size: S"x30"
sec.): 69.4 .2
4 4
3
Pres.(stnd otm.J:20.5
Ye'TI
Dole: 2-26-32
Airfoi/.-NA.CA. 2RZ 12 R.N:3,130,000 -l2
w
Where
tested-L.R.A.L.
Test:
l!D.T.
765
Date: 2-26-32
-8 u
Test: V. D. 765 _/6
0 -.4
Corrected for tunnel-wol/effect -.4
Corrected to inf/Hite aspect ratio
-8 -4 0 4 8 12 16 20 24 28 32
04 -2 0 .2 .4 .6 .8 /.0 1.2 /.4 /.6 1.8
Angle of atlock, a (degrees)
Lift coefficient,C
Flovaa 79,-N.A. C.A. 2Rt12 airfoil.
40 7.63 -a96
l2
"
45
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 48
0 0 0
L251.9 -1.99
25 261 -261
50 .55 -3.55
10 20 -4.20
/0469 - 4.69
0
0
2.61 -26/
3.55 -3.55
4.20 -4.20
205.74 -574
5.74-5.74
/55.35-5.35
o+x NA.C.A. 0012E
0 00
0012E
/.89-/.89 10
20 40 60 BO
Per centofcha'd
O
4.68-4.69
5.35 -535
255.94 -594 594 -594
30.00 -.00 .00 -6.00
40 .74 -55.74 5.74-5. 74
50 .3/ -4.3/ 4.31-4.31
60 /.96 -/.96 /.96-/96
70 .So - .50 .50-50
90.50- .50- .4/ -/.43
90 .50 - .50 -3.23 4.30
95 5 - .50 -45-6.5
5.
7
0
p
t
O
C
se i-
98.79
0011501(-.591
GO
O E.Rnd. ^.5e
28 0
240 20
^
C 1
c.p.
_
C.P .
20 40
/660
0 40
0 4m
-4 0.
.D:
360)I
2.0
.40
L1$
-.01
3z o
/.6
.36
.0;
/.6
.32
0.01
U
Q
W
u
-4
?0 t
162
m
4
8
/2 /6
20 24 28 32
Angle of ottock. a (degrees)
/z ^c
o!o
.0;
a
.4 .08
.04
C
\
0 4:_
0 •^
NA.CA.00120 8 0012F R. N: 3,230.000
5Ve%(ft/sec.):
"x30"
68.1
U
Pres.(sti7dotm.):210 Dote:7(V 14-32
`^c
Where tested. L.MA.L. VD.T. 7519753
Corrected for tunnel-wall effect
-•'
0
X8.0'
'4 0
/. a.24v ro.0,
vk
o a
.6 ./2
e
04
oe
/4 .28-- ^.0:
.8 v.16o
6 0/00
I0
/(
.44
.20
o 12 ^ 80--t-
$4
100
4,Y
u
00
'' C
N. AC 002 a80012F, I R.N.:3,230,000
Da1e: 1-(9812)-32 Test: I D.T 751 8 753
Corrected to infinite ospectmtio
4t
v
8 ^'
./2e
/6
0 .2 .4 .6 .8 /.0 1.2 /.4 /.6 /.8 ZU
Lift coefficient, C
FIGURE 80 .-N.A.C.A. 0012Fo and 0012F, aufoas.
PRECISION
largely eliminated during the later tests. For this
A general discussion of the errors and corrections
involved in airfoil testing in the variable-density tunnel
is included in reference 8. In connection with this
report, it was hoped that a more specific discussion of
report, however, the effect of the error from this source
has been minimized by repeating the tests of many of
the airfoils, including all of the symmetrical series
the various sources of error and separate estimates of
the various errors might be given. However, after a
careful study of all the measurements it became
apparent that practically all the errors may be regarded
as accidental; that is, of the type the magnitude of
which may best be estimated from the dispersion of the
results of independent repeat measurements. The
major portion of these errors is caused by insufficient
sensitivity of the balance and manometers, by the
originally reported in reference 2.
The magnitude of all such accidental errors was
judged from the results of repeat tests of many
airfoils, and from the results of approximately 25
tests of one airfoil that were made periodically throughout the investigation to check the consistency of the
measurements. The accidental errors in the results
presented in this report are believed to be within the
limits indicated in the following table:
Cm,
CD1(CL=0)
V
t 0.15o
0.01
- 0.03
f 0.003
0.0006
1 -0.0002
CD0 ( CL -1)
0.0015
- 0.0008
a
personal error involved in reading mean values of
slightly fluctuating quantities, and by the error due to
slight surface imperfections in the model. The last is
perhaps the most serious source of error. The models
were carefully finished before each test, but the presence of particles of hard foreign matter in the air stream
tended to cause a slight pitting of the leading edge of
the model during each test. This pitting was probably
the major source of error in connection with the earlier
tests, but it was reduced for the later tests when the
necessity of a more careful inspection of each model
was appreciated. After a considerable period of
running the particles in the tunnel were found to become lodged, permitting this source of error to be
CLmax
In addition to the consideration of the accidental
errors, all measurements were carefully analyzed to
consider possible sources of errors of the type that
would not be apparent from the dispersion of the
results of repeat tests. A rather large (approximately
1.5 percent) error of this type is present in all the air-
46
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
velocity measurements resulting from a reduction in
the apparent weight of the manometer liquid when the
For the purpose of comparing the results from different wind tunnels and of applying these results to air-
density of the air in the tunnel is raised to that cor-
planes in flight, it is also necessary to consider the
responding to a pressure of 20 atmospheres. The
effects of this error, however, are reduced by the presence of another error in the air-velocity measurements
due to the blocking effects of the model in the tunnel.
The measured coefficients, obtained by dividing the
measured forces by %pV 2 , as well as the derived coefficients are, of course, affected by errors in the airvelocity measurement. Aside from this source of
effects of air-stream turbulence. In air streams having
error, it is believed that only two other sources need
be considered: first, the deflection of the model and
supports under the air load; and second, the interference of the airfoil supports on the airfoil. The
angle of attack and the moment coefficient are affected
by the deflection of the airfoil and supports. The
error in angle of attack, which is proportional to
Cmcl,,
was found to be approximately —0.1° for an
different degrees of turbulence, the value of the
Reynolds Number cannot be considered as a sufficient
measure of the effective dynamic scale of the flow
The airfoil characteristics presented in this report were
obtained at a value of the Reynolds Number of approximately 3,000,000, which corresponds roughly to the
Reynolds Number attained in flight by a mediumsized airplane flying near its stalling speed. Consideration of the effects of the turbulence present in the
variable-density tunnel (see references 11 and 12) leads,
however, to the belief that these results are more
nearly directly applicable to the characteristics that
would be obtained in flight at larger values of the
Reynolds Number.
N./2
tSymmetrii al ai^R!
foi i I
,ar
oma
I 27 per rodlan
x 4%
04
Vy
+ 6%
ro
°
Maximum thickness in per cent of chord
0
Camber position in fraction of chord
(Abscissa of maximum mean-line ordinate)
F1aunE 81.—Variation of liftcurve slope with thickness.
airfoil having a moment coefficient of —0.075. The
error from this source in the moment coefficient is
inappreciable at zero lift, but at a lift coefficient of 1
may amount to —0.001. The errors resulting from
the support interference are more difficult to evaluate,
but tests of airfoils with different support arrangements
lead to the belief that they are within the limits indicated in the following table:
«
±0.05°
0.00
— 0.02
CLmax
Cmcl4
CDa(CL=0)
CDa (CZ =1)
10.001
f 0.0002
1
0.0000
±0.0010
FIGURE 82.—Variation
of lifGcurve slope with camber. Results for 12 percent thick
airfoils.
DISCUSSION
The results of this investigation are here discussed
and analyzed to indicate the variation of the aerodynamic characteristics with variations in thickness
and in mean-line form. For the analysis of the effect
of thickness, test data from consecutive tests of airfoils
having different thicknesses and the same mean-line
form are used. The analysis of the effect of the meanline form is made with respect to consecutive tests of
airfoils of the same thickness (12 percent of the chord)
and related mean-line forms. The results are compared, where possible, with the results predicted by
thin-airfoil theory, a summary of which is presented
The tunnel-wall and induced-drag corrections applied to obtain the airfoil section characteristics might
in the appendix.
also be treated as sources of systematic errors. Such
Lift curve.—In the usual working range of an airfoil section the lift coefficient may be expressed as a
errors need not be considered, however, if the section
characteristics are defined as the measured characteristics with certain calculated corrections applied.
Errors in the tunnel-wall corrections, however, should
be considered when the results from different wind
tunnels are compared. For consideration of these
errors, the reader is referred to references 9 and 10.
LIFT
linear function of the angle of attack
CL =ao (ao — aaa)
where a n is the slope of the lift curve for the wing of
infinite aspect ratio and a Le is the angle of attack at
zero lift.
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL
47
The variation of the lift-curve slope with thickness
is shown in figure 81. The points on the figure rep-
given mean line without altering the camber position.
resent the deduced slopes as measured in the angular
as the position of the camber moves back along the
chord. The experimental values are compared with
range of low profile drag. These results confirm
previous results (reference 1) in that they show the
lift-curve slope to decrease with increasing thickness.
The theory also predicts an increased negative angle
the theoretical values in figures 83 and 84. The ex-
The camber has very little effect on the slope, as
indicated in figure 82, although a rearward movement
of the position of the camber tends to decrease the
lot
003
0 43
024
0 25
0 44
0 45
o63
0 64
0 65
a.9t
u, X
63
E v,
a
v o
av
v^
`64
g14.&
y
a,
v
i
65
Comber position in froction of chord
fAbscisso of maximum --line ordinate)
704
FIGURE 83.—Variation of angle of zero lift with camber. Points shown are for 12
percent thick airfoils. Curves indicate general trends for the different thick-
8
12
16
20
Maximum thickness in percent of chord
24
Francs 84.—Variation of angle of zero lift with thickness. Numbers refer to mean-
nesses.
camber designation.
/.f
/. E
I
/.5
b
/.e
/.t
CL_
.t
° 'Symmefrfcol series
2% mean camber
0 23 series
••
x 24
+ 25
04
4%mean camber
o 43 series
••
x 44
+ 45
6Z mean
6 /2 16 20 4 8 /2 16 20 4 8 e 16 20 24
Maximum thickness in percent of chord
FmURE
85.—Variation of maximum lift with thickness.
slope slightly. Table II gives the numerical values of
the slope in convenient form for noting the general
trends with respect to variations in thickness and in
camber.It will be noted that all values of the slope
lie below the approximate theoretical value for thin
wings, 2a per radian; the measured values lie between
95 and 81 percent, approximately, of the theoretical.
The angle of zero lift is best analyzed by means of a
comparison with that predicted by the theory. Thinairfoil theory states that the angle of zero lift is pro-
portional to the camber if the camber is varied, as
with these related airfoils, by scaling the ordinates of a
perimental values lie between 100 and 75 percent,
approximately, of the theoretical values, the departure becoming greater with a rearward movement of
the position of the camber and with increased thickness
(above 9 to 12 percent of the chord). Numerical
values of the angle of zero lift are given in table III.
Maximum lift.—The variation of the maximum lift
coefficient with thickness is shown in figure 85. It
will be noted that the highest values are obtained with
moderately thick sections (9 to 12 percent of the chord
thick, except for the symmetrical sections for which
the highest values are obtained with somewhat thicker
s
REPORT NATIONAL ADVISORY
48
sections). The variation with camber, shown in
figure 86, confirms the expected increase in maximum
lift with camber. The gain is small, however, for the
normal positions of the camber, but becomes larger
as the camber moves either rearward or forward. It
will be seen by reference to figure 85 that the camber
becomes less effective as the thickness is increased.
This reduced effectiveness of the camber is in agreement with a conclusion reached in reference 13 that
;OMMITTEE FOR AERONAUTICS
bar (for airfoils having normal camber positions;
i.e., 0.3c to 0.5c).
MOMENT
Thin-airfoil theory separates the air forces acting on
any airfoil into two parts: First, the forces that produce a couple but no lift (they are dependent only on
the shape of the mean line); second, the forces that
produce the lift only, the resultant of which acts at
for airfoils having a thickness ratio of approximately
Maximum thickness in per cent •,
-.05
20 percent of the chord, camber is of questionable
i
value. Numerical values of the maximum lift co-
69
j.
' L5
-.04
efficient are given in table IV.
Air-flow discontinuities.—These and other windtunnel tests indicate that at the attitude of maximum
Theoretico/ curvy
2/
w -.03
^E
U
o -.02
-.0/
_L
4
Comber position in fraction of chord
(Abscissa of maximum mean-tine ordinate)
FIGURE 87.—Variation of moment at zero lift with Camber. Points shown an for
12 percent thick airfoils. Craves Indicate general trends for the different thick-
nasses.
.90
Ca„
0 23 — 0 43
0 63
o25
0 65
024 _1 0 44
o 45
0 64
o ,.BL
E m
Qt
V$
C• 74
i
_L
4
Comber position to rrau/on or enora
(Abscissa of maximum mean-line ordinate)
FIGURE 88. —Variation
of maximum lift with camber. Results for 12 percent thick
airfoils.
lift the air forces on certain airfoils exhibit sudden
changes which in many instances result in a serious
loss of lift. The probable cause of these air-flow discontinuities is discussed briefly in reference 13. The
stability or instability of the air flow at maximum
4
20
16
8
12.
Maximum thickness in percent of chord
a4
FIGURE 88.—Variation of moment at zero lift with thickness. Numbers refer to
mean-camber designation.
a fixed point. We then have in the working range an
expression for the total moment taken about any
point
Cm = C,ea+nCL
are classified into three general types as noted in
table IV, but the degree of stability is difficult to
judge. It may be generally concluded that improved
stability may be obtained by (1) having a small
where Co is the moment coefficient at zero lift and
nCL is the additional moment due to lift.
As with the angle of zero lift, the theory states that
the moment at zero lift is proportional to the camber
and predicts an increase in the magnitude of the
moment as the camber moves back along the chord.
leading-edge radius, which causes an early break-
Figures 87 and 88 show the values of the moment
down of the flow with a consequent low value of the
maximum lift, (2) increasing the thickness (beyond the
normal thickness ratios), or (3) increasing the cam-
coefficient as affected by variations of camber and
lift may be judged by the character of the lift-curve
peaks indicated for the various airfoils. The curves
thickness compared with the theoretical values. Referring to figure 87, the plotted data indicate that the
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL moment coefficients are nearly proportional to the
camber. It will also be noted that the curves representing the ratios of the experimental coefficients to
the camber are nearly parallel to the equivalent curve
representing the theoretical ratios except that the
curves tend to diverge for positions of the camber
well back. Figure 88 shows that the experimental
values lie between 87 and 64 percent, approximately,
of the theoretical. Numerical values of the moment
coefficient at zero lift are given in table V.
the profile drag as the minim um value plus an additional drag dependent upon the attitude of the airfoil,
we have in coefficient form
CD = CD,+ (CDamtn + ACDa)
The induced-drag coefficient C U, which is computed
by means of the formula given in reference 8, is considered to be independent of the airfoil section. The
variation of the profile-drag coefficient with the shape
variables of the airfoil section is analyzed with respect
0 Ox mean combe .
x 2%
a 4%
+ 6%
.04
.02
49
Symmetric./ oirfoixe
*
C&-
q
+0.
Norma/cambered oirfoi/s
+
4
8
0
16
20
24
Maximum thickness in per cent of chord
FIGURE 89.—Variation
of position of constant moment with thickness. Values of
a for equation .,,, .s +nC . Results for airfoil. having normal camber
positions (0.3c to 0.6c).
If the resultant of the lift forces acted exactly
through the quarter-chord point, as predicted by the
theory of thin airfoils, there would be no additional
moment due to the lift when the moments are taken
about this point. The curves of C ,, against CL,
however, show a slope in the working range which
indicates that the axis of constant moment is displaced
somewhat from the quarter-chord point. The factor n
represents the amount of this displacement as obtained
from the deduced slopes of the moment curves in the
Mean camber
)2
O
4
s%
4%
Symmetrical airfoil
zr,
4
.6
.8
Comber position in fraction of chord
(Abscissa of maximum mean-line ordinate)
.2
in fraction
FIGURE 91.—Varlation
of minimum profile dmg with thioknom far the symmetrical
airfoils.
to the variations of the two components of the profile
drag.
Minimum profile drag.—The variation of the minimum profile-drag coefficient with thickness for the
symmetrical sections is shown in figure 91. The cambered sections show the same general variation with
thickness but, to avoid-confusion, the results are not
plotted. The variation of the minimum profile-drag
.00&
Mean ca, bar,
V
.Co a
4X
2%
/.O
names 90.—Variation of position of constant moment with camber. Values of n
fmnquationC..l,—C.,+nCz. Results for 12 percent thick airfoils.
normal working range. The variation of this displacement with thickness and with camber is shown in
figures 89 and 90. Table VI gives the numerical
values. Beyond the stall all the airfoils show a sharp
increase in the magnitude of the pitching moment.
The suddenness of this increase follows the degree of
stability' at the stall as indicated by the type of the
lift-curve peak.
DRAG
The total drag of an airfoil is considered as made up
of the induced drag and the profile drag. Considering
O
s
.8
.4
.2
Comber position
in fraction of chord
(Abscissa of
LL
mu. mean-11ne ordinate)
92.—Incmase in minimum profiledmg due to camber. Reseltsforl2pereent
thick airfoils. Values of k for equation Cx,... k+0.0060+0.01 9+01 V, where k
is the increase in CDs.., due to camber and t is the maximum thickness in fraction
of chord.
FIGURE
coefficient with the profile thickness may be expressed
by the empirical relation
CDamfn =k +0.0056+0.01t +0.112
where t is the thickness ratio and k (which is approxiinately constant for sections having the same mean
Line) represents the increase in C,,,.,. above that
computed for the symmetrical section of corresponding
thickness. The variation of CDanfa with camber is
indicated by the variation of k as shown in figure 92.
50
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
The effect of camber is small except for the highly
cambered sections having the maximum camber well
back. Numerical values of CDamr° are given in table VII.
Additional proflle drag.—The additional profile
drag, which is dependent upon the attitude of the airfoil, has previously been expressed as a function of the
lift (reference 4) by the equation
ACDe = CDO — CDOmtn = 0.0062 (CL — C,,.,,)'
where CL °n, may be called the optimum lift coefficient;
that is, the lift coefficient corresponding to the minimum profile-drag coefficient. This equation holds
.028
.024
0
x
n
.020
J
This function is represented in figure 93 as the curve
determined from the results for the symmetrical airfoils and for the airfoils having a camber of 2 percent
of the chord. As the camber is increased, the dispersion of the plotted points from the curve becomes
greater. In general the points above the curve correspond to thick sections and sections in which the maximum camber is well back. The departure from the
curve becomes greater with increased thickness and
with a rearward movement of the maximum-camber
position. The points well below the curve correspond
to the thin airfoils.
OX mean comber
2
4 % .,
6%
x
I°
—
°-1
016
xX#•
11.012
.008
.004
:° tx
.2
.4
.3
.5
a-a-,
.6
.7
.8
.9
40
F—E 93.—Addaioml pmdle drag.
reasonably well for the normally shaped airfoils at
values of the lift coefficient below unity.
A convenient practical method of allowing for the
increased values of CD, at moderately high values of
the lift coefficient is to include the additional profile
drag with the induced drag, as suggested in reference
2. For the symmetrical airfoils of moderate thickness
the term to be added to the induced-drag coefficient
was given as 0.0062 CLE . The relative importance of
this term may be better appreciated by considering
that it represents 11.7 percent of the induced drag of
an elliptical airfoil of aspect ratio 6. The same
method may also be applied to other airfoils if the
value of the optimum lift is not too large.
Andrews (reference 14), using the part of these data
published in references 2, 4, and 5, suggests for the
additional profile drag the form
A
D0
=f1
ODr
CL— CL
_%L,
CL
1
Because the additional profile drag is not a simple
function of the lift, and also because the results as
presented in figure 93 are difficult to follow, generalized
curves for the relation
OCDe = f
(CL— CLavr)
are given in figure 94. These curves are given to
represent more accurately the additional profile drag
for the normally shaped sections.
Optimum lift.—The optimum lift, as defined above,
is the value of the lift corresponding to the minimum
profile drag. As the determination of this value of the
lift is largely dependent upon the fairing of the profiledrag curves, special curves were faired for this purpose
on enlarged-scale plots corresponding to certain related
airfoils grouped together. The values of the optimum
lift coefficients obtained in this manner are given in
table VIII. It may be noted by reference to this table
that the optimum lift coefficient increases with camber
CHARACTERISTICS OF AIRFOIL SECTIONS FRO] I TESTS IN VARIABLE-DENSITY WIND TUNNEL
and for the highly cambered sections a definite increase
accompanies a forward movement of the camber.
.02
2..16.
"eon
Max, lhickness 69. 4.
0 07. comber
v,*
0
.02
.o/
I
29,—••
x
ma
p
u
.0/
o 49,
+ 6%—
I.
" -T1
1 1
1
6
IT
in table VIII representing the theoretical lift coefficients at the "ideal" angle of attack for the mean
line; i.e., the angle of attack for which the thin-airfoil
theory gives a finite velocity at the nose. (See the
appendix.)
410'2-
GENERAL EFFICIENCY
Maximum lhickness 9%
The general efficiency of an airfoil cannot be expressed by means of a single number. The ratio of
0
0
`6 4° 2 0°
.02
51
it is not primarily dependent upon the shape of the
mean line. Nevertheless it is interesting to compare
the optimum lift coefficients with the values included
Symmetrical airfoil
20
k
Maximum lhickness /2%
.0/
r
I
Vg
^
E
0
4
6
x$.02
Meon rnmber
f)-
2
1
^^
Maximum lhickness /5%
G°
.0/
O
ov
^F
U
U
a.rA ` . er•
.02
J B 1
O,
6
4 i
Maximum thickness /B9,
ow
.0/
$
^o
u
0 .oFmo.
1 4° 2 O•.
.02
Comber powhbg in fraction of chord
(Abscissa of maximum mean-tine ordinate/
Maximum thickness 219,
.Ol
O
FIGURE 90.—Variation of CI,_lC o.c. witb camber. Results arc tot 12 pereeut
thick airfoils.
O
2
.4
.6
.8
10
/.2
14
FIGURE 91.—Additional
profile dmg as a function of Cy— C.,1. Results era for sitfalls having normal camber positions (0.5c to 0.5c).
the maximum lift to the minimum profile drag is, however, of some value as the measure of the efficiency of
an airfoil section. The variation of this ratio with
thickness is shown in figure 95. The curves of this
figure indicate that the highest values of the ratio are
Symmetrical
series
200
160
V^/20
v 80
40
comber
0 23 series
_j2%mean
x 24
+25
4%mean comber
o 43 series
4
x 44
6%mean comber
o 63 series
x 64
+65
+
4 8 12 16 20 4 6 /2 /6 20 4 8 12 16 20 24
Maximum thickness in percent of chord
0
FIGURE 95.—Veriatfou of C ao,/CD,. with thickness.
More important than these variations, however, is the given by the sections between 9 and 12 percent of the
variation with thickness. The rapid decrease in the chord thick. The variation with camber, shown in
optimum lift with increased thickness indicates that figure 96, is less important. An increase in the camber
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
52
above 2 percent of the chord and a rearward move-
ment of the camber (for the highly cambered sections)
tend to decrease the value of CL,aee/CDO,afa. The
numerical values of the ratio are given in table IX.
seetlon
T series
Normal
11-11
11
0006
0012
0018
0.10
.40
.89
0.40
1.58
3.56
1.19
3.80
7.15
SUPPLEMENTARY AIRFOILS
For the purpose of investigating briefly the effects
of certain shape variables other than those discussed
in the main body of the report, 10 supplementary air-
foils were tested. The airfoil sections were as follows:
6 symmetrical sections with modified nose shapes, 2
sections with reflexed mean lines, and 2 sections simulating those of a wing having a flexible trailing edge.
Maximum thickness 0.12c
/.4
^Moximum thickness 0.18c
/.2
/.0
are plotted against the leading-edge rad ii in figure 97.
It is interesting to note that the leading-edge radius is
very critical in its effect on the maximum lift when
the radius is small. This critical effect is also indicated
by the rapid increase in the maximum lift with increasing thickness for the thin sections as shown in figure 85.
Airfoils with reflexed mean lines.-Previous investigations have shown that the pitching moment of
cambered airfoils can be reduced by altering the form
of the mean line toward the trailing edge, with a consequent loss of maximum lift but only a small reduction
in drag. In order to compare the characteristics of
sections of this type with those of the related sections
Moximum thickness 006e
of normal form, two airfoils were developed with the
basic thickness distribution of the N.A.C.A. 0012 disposed about certain mean lines of the form given in
reference 15
y^=hx(1-x) (1-Tx)
.6
.4
.2
O
The aerodynamic characteristics of the modified
sections are given in figures 72 to 77. These may be
compared with the characteristics of the normal sections given in figures 4, 6, and 8. The maximum lift
coefficients of the modified and the normal sections
/
2
3
4
5
6
i-Hnn-eons radus in Der cent of chord
F-RE
7
97.-Variation of maximum lift with nose radius.
Airfoils with modified nose shapes.-The airfoils of
the first supplementary group investigated were developed from three of the symmetrical N.A.C.A. family
airfoils: The N.A.C.A. 0006, the N.A.C.A. 0012, and
the N.A.C.A. 0018. For each of these basic (or normal) sections one thinner-nosed section, denoted by
the suffix T, and one blunter-nosed section, denoted
by the suffix B, were developed and tested. The
derivation of each modified section was similar to that
of the normal section and was accomplished by a systematic change in the equation that defines the normal
section. This change is principally a change in the
nose radius, but it also results in modifications to the
profile throughout its length, except at the maximum
ordinate and at the trailing edge. The nose radii of the
sections in percent of the chord are as follows:
The values of h in this equation were chosen to give a
camber of 0.02 and the values of X were chosen to give
the airfoil designated the N.A.C.A. 2R 112 a small
negative moment and the airfoil designated the
N.A.C.A. 2R 212 a small positive moment. Characteristic curves for the two airfoils are given in figures
78 and 79. The principal characteristics of the sec-
tions may be conveniently compared with those of the
related symmetrical section, the N.A.C.A. 0012, and
a related normal section having a camber of 2 percent
of the chord, the N.A.C.A. 2412, by means of the
following table arranged in the order of increasing
pitching-moment coefficients.
c.
C.p
section
Cam„
211M
1.47
0.0086
171
0.004
0012
2R^12
2912
1.53
1.63
1.02
.0083
.0083
.0085
184
184
IN
-.002
-.020
-.044
CDC- CDC-
These results indicate that airfoils having reflexed
mean lines may be of questionable value because of the
adverse effect of this mean-line shape on the maximum
lift coefficient.
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL Thickness and camber modifications near the
trailing edge.—Two airfoils were developed to simulate an airfoil having a flexible trailing edge in a straight
and in a given deflected position. The thickness distribution is composed of three parts: the forward portion (0 to 0.3c) having the same distribution as the
N.A.C.A. 0012, the rear portion (from 0.7c to the
trailing edge) having a thin, uniform value, and the
central portion joining these two with fair curves.
As shown in figure 80, the two airfoils differ only in
the rear portion, the section designated N.A.C.A.
0012Fo simulating that of a wing having the trailing
edge deformed for the high-speed condition, and the
section designated N.A.C.A. 0012F I simulating that
of the same wing with the trailing edge bent down in a
circular arc. Curves of the aerodynamic characteristics for both conditions are compared in figure 80.
Considering the results given by both airfoils as two
conditions for one airfoil, a very high maximum lift
with a reasonably low minimum drag is obtained.
On this assumption the ratio
CL'C."ef. Is
197, slightly
higher than the value of this ratio given by the
N.A.C.A. 2412.
In order to study the effects of an extreme change in
the thickness distribution, the principal characteristics of the two sections may be compared with those
of the related normal sections, the N.A.C.A. 0012 and
the N.A.C.A. 6712. The maximum lift coefficient is
little affected by the change in the thickness distribution, but it is of interest to note (table I) that the slope
of the lift curve of the N.A.C.A. 0012Fa is slightly
greater than 2a per radian, as compared with an appreciably lower slope for the N.A.C.A. 0012. The profile
drag is also affected by the change in the thickness
distribution. Of the two symmetrical sections, the
profile drag of the N.A.C.A. 0012Fp is much higher
than that of the N.A.C.A. 0012 over the entire lift
range. This is not true, however, for the two cambered sections. Comparing the characteristics of the
N.A.C.A. 0012F, with those of the N.A.C.A. 6712, we
find that at low values of the lift the profile drag of
the former is much higher, but as the lift increases this
difference becomes less, and in the high-lift range the
profile drag of the N.A.C.A. 0012F, is considerably
less than that of the N.A.C.A. 6712.
CONCLUSIONS
The variation of the aerodynamic characteristics of
the related airfoils with the geometric characteristics
investigated may be summarized as follows:
Variation with thickness ratio:
1. The slope of the lift curve in the normal working
range decreases with increased thickness, varying
from 95 to 81 percent, approximately, of the theoretical
slope for thin airfoils (2v per radian).
53
2. The angle of zero lift moves toward zero with
increased thickness (above 9 to 12 percent of the
chord thickness ratios).
3. The highest values of the maximum lift are obtained with sections of normal thickness ratios (9 to
15 percent).
4. The greatest instability of the air flow at maximum lift is encountered with the moderately thick,
low-cambered sections.
5. The magnitude of the moment at zero lift decreases with increased thickness, varying from 87 to
64 percent, approximately (for normally shaped airfoils), of the values obtained by thin-airfoil theory.
6. The axis of constant moment usually passes
slightly forward of the quarter-chord point, the displacement increasing with increased thickness.
7. The minimum profile drag varies with thickness
approximately in accordance with the expression
C-i.,n =k+0.0056 +0.0lt+0.lt2
where the value of k depends upon'the camber and t is
the ratio of the maximum thickness to the chord.
8. The optimum lift coefficient (the lift coefficient
corresponding to the minim um profile-drag coefficient)
approaches zero as the thickness is increased.
9. The ratio of the maximum lift to the minimum
profile drag is highest for airfoils of medium thickness
ratios (9 to 12 percent).
Variation with camber:
1. The slope of the lift curve in the normal working
range is little affected by the camber; a slight decrease
in the slope is indicated as the position of the camber
moves back.
2. The angle of zero lift is between 100 and 75 percent, approximately, of the value given by thin-airfoil theory, the smaller departures being for airfoils
with the normal camber positions.
3. The maximum lift increases with increased camber, the increase being more rapid as the camber
moves forward or back from a point near the 0.3c
position.
4. Greater stability of the air flow at maximum lift
is obtained with increased camber if the camber is in
the normal positions (0.3c to 0.5c).
5. The moment at zero lift is nearly proportional to
the camber. For any given thickness, the difference
between the experimental value of the constant of
proportionality and the value predicted by thin-airfoil
theory is not appreciably affected by the position of
the camber except for the sections having the maximum
camber well back, where the difference becomes
slightly greater.
6. The axis of constant moment moves forward as
the camber moves back.
7. The minim um profile drag increases with increased camber, and also with a rearward movement
of the camber.
54
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
8. The optimum lift cofficient increases with the
camber and for the highly cambered sections a definite
increase accompanies a forward movement of the
camber.
9. The ratio of the maximum lift to the minimum
profile drag tends to decrease with increased camber
(above 2 percent of the chord) and with a rearward
movement of the camber (for the highly cambered
sections).
LANGLEY MEMORIAL AERONAUTICAL LABRATORY,
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS,
LANGLEY FIELD, VA., December 20, 1932.
APPENDIX
It is proposed in this section of thereport to present,
briefly, a summary of the results of the existing thinairfoil theory (based on the section mean line) as applied to the prediction of certain section characteristics. Such a summary is desirable because at present
the results must be obtained from several different
sources which give them in a form not easily applied.
Three characteristics are considered; namely, (1) the
angle of zero lift aLO, (2) the pitching-moment coefficient C d4, and (3) the "ideal" angle of attack a,,
or the corresponding lift coefficient Czl, that is, values
corresponding to the unique condition for which the
theory gives a finite velocity at the nose of the airfoil.
(See reference 16.)
Expressions for lift and moment coefficients may be
written as follows if the angles are measured in radians:
Cy=27r(a-%) (1)
CLl=27r(aj-a4)
(2)
Gme/4 = 2( 13+aL0)
(3)
If the leading end of the mean line is chosen as the
origin of coordinates and the trailing end is taken on
the xaxis at x=1, then the parameters aLo, a,, and
j3 are given by the following integrals
azo =
fly fl (x) dx
(4)
0
ar =
f y f2(x) dx
(5)
4(1-2x)
f3(x)-7[x(1-x)]11
and y is the ordinate of the mean line at a given abscissa
x. The integrals (4) and (6) may be shown to be
identical with the corresponding integrals given by
Glauert (reference 15) and by Munk (reference 17),
and integral (5) is given by Theodorsen (reference 16).
The evaluation of these integrals for the N.A.C.A.
airfoil sections given in this report was accomplished
analytically. The values of aL, (changed from
radians to degrees), C-c4 and CL„ so computed, are
given in tables III, V, and VIII, respectively, in the
main body of the report. This method of evaluation,
however, cannot be applied to many of the commonly
used sections because they do not have analytically
defined mean lines; hence, an approximate method
must be used. A graphical determination gives good
results and for convenience the values of the three
functions, (7), (8), and (9), at several values of x, are
given in the following table:
x
0
0.0125
.0260
.0500
.0750
.1000
.15
.20
.25
f1W
x
111.17
7.747
5.258
9.109
3.395
2.496
1.910
1.470
0.30
.40
.60
.60
.70
.80
.90
.95
1.00
f,(-)
W4
W4
-0.992
1.111
-1.083 0.662
-1.273 0.271 0.520
-1.624 -.271 -.620
-2.315 -.662 -1.111
-3.978 -1.492 -1.910
-10.61 -4.718 -3.395
-29.21 -13.84 -5.258
In general, some difficulty would be expected with
the graphical method because the values of the above
functions tend to infinity at the leading and trailing
edges. Actually, because the ordinates of the meanline extremities are zero, the integrand may approach
zero, and does at the leading edge for the integral (4),
and at the leading and trailing edges for the integral (6).
Difficulty, however, is encountered at the trailing edge
for the integral (4) and at the leading and trailing edges
for the integral (5). In order to avoid this difficulty,
integral (4) is evaluated graphically from x=0 to
x=0.95, and the increment contributed by the portion from x=0.95 to x=1 is determined analytically.
Likewise, integral (5) is evaluated graphically from
x=0.05 to x=0.95 and analytically for the extremities.
The analytical determination of the increments is
accomplished by assuming the mean line near the
ends to be of the form
(6)
Evaluating the integrals gives
(7)
daze = -0.964yo .s5 +0.0954y;
(x=0.95 to x=1)
Aa,= +0.467y,.05 +0.0472yo
(x=0 to x=0.05)
-0.467yo. 95 +0.0472y1
(x=0.95 to x=1)
where yo and yf are the mean-line slopes at the
leading and trailing edges, respectively.
where
-1
fl(x)= 7r(1-x) [x(1-x)]E
(1- 2x)
f2(x)=2,'[x(1-x)]n
f3W
y=a+bx+cc2
t
d= f
0
r,(x)
-2.901 118.15
-2.091 39.73
-1.537 13.84
-1.308 7.403
-1.178 4.716
-1.048 2.447
2.995 1.492
-.9so .980
0
yf3(x) dx
(9)
(8)
I
CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL
REFERENCES
1. Jacobs, Eastman N., and Anderson, Raymond F.: LargeScale Aerodynamic Characteristics of Airfoils as Tested
in the Variable-Density Wind Tunnel. T.R. No. 352,
N.A.C.A., 1930.
2. Jacobs, Eastman N.: Tests of Six Symmetrical Airfoils in
the Variable-Density Wind Tunnel. T.N. No. 385,
N.A.C.A., 1931.
3. Pinkerton, Robert M.: Effect of Nose Shape on the Characteristics of Symmetrical Airfoils. T.N. No. 386,
N.A.C.A., 1931.
4. Jacobs, Eastman N., and Pinkerton, Robert M.: Tests of
N.A.C.A. Airfoils in the Variable-Density Wind Tunnel.
Series 43 and 63. T.N. No. 391, N.A.C.A., 1931.
5. Jacobs, Eastman N., and Pinkerton, Robert M.: Tests of
N.A.C.A. Airfoils in the Variable-Density Wind Tunnel.
Series 45 and 65. T.N. No. 392, N.A.C.A., 1931.
6. Jacobs, Eastman N., and Pinkerton, Robert M.: Tests of
N.A.C.A. Airfoils in the Variable-Density Wind Tunnel.
Series 44 and 64. T.N. No. 401, N.A.C.A., 1931.
7. Jacobs, Eastman N., and Ward, Kenneth E.: Tests of
N.A.C.A. Airfoils in the Variable-Density Wind Tunnel.
Series 24. T.N. No. 404, N.A.C.A., 1932.
S. Jacobs, Eastman N., and Abbott, Ira H.: The N.A.C.A.
Variable-Density Wind Tunnel. T.R. No.416, N.A.C.A.,
1932.
55
9. Higgins, George J.: The Prediction of Airfoil Characteristics. T.R. No. 312, N.A.C.A., 1929.
10. Knight, Montgomery, and Harris, Thomas A.: Experimental Determination of Jet Boundary,Corrections for
Airfoil Tests in Four Open Wind Tunnel Jets of Different
Shapes. T.R. No. 361, N.A.C.A., 1930.
11. Stack, John: Tests in the Variable-Density Wind Tunnel to
Investigate the Effects of Scale and Turbulence on Airfoil Characteristics. T.N. No. 364, N.A.C.A., 1931.
12. Dryden, H. L., and Kuethe, A. M.: Effect of Turbulence in
Wind Tunnel Measurements. T.R. No. 342, N.A.C.A.,
1930.
13. Jacobs, Eastman N.: The Aerodynamic Characteristics of
Eight Very Thick Airfoils from Tests in the VariableDensity Wind Tunnel. T.R. No. 391, N.A.C.A., 1931.
14. Andrews, W. R.: The Estimation of Profile Drag. Flight,
vol. XXIV, no. 25, pp. 530a-530d, 1932 and no. 31, pp.
710a-710c, 1932.
15. Glauert, H.: The Elements of Aerofoil and Aircrew Theory.
Cambridge University Press (London), 1926.
16. Theodorsen, Theodore: On the Theory of Wing Sections
with Particular Reference to the Lift Distribution.
T.R. No. 383, N.A.C.A., 1931.
17. Munk, Max M.: Elements of the Wing Section Theory and
of the Wing Theorv. T.R. No. 191, N.A.C.A., 1924.
c,
A
HIR
MI i
E.
H M999
as 2vM222
.....
ttSS2 22
...........
9
.2 -2
2i 9 !-f E i R
t3 ^3
. -9 .2 .2 -S
2
iR iR v t S a v tg
t.2 .
2
-!2 -2-2.2
e
8
a! Sg !2 8;
ii i2vi & N :,.^ tw
N
v v
.. ....
Rns4 :-. 2 age
4 2:^ & 4 k v Ni i v 4 8 z
52
.....................
................
s
-8 115
tgma
-S -9 ...... ;
. .... .. .
-2 -e -2
4aa3^5tl 2 i^k^t-,st al
u a
- -----
-S -2 -S -S! -2 -2 -S -S
I- ri I rri I I I I I I rl I I I I rl I I I I I I I I I I I
tngR
sees
eva 2 a 8 t aa
...............
?a r-.r -.-.8 v
^cl
cp
CHARACTERISTIC S OF AIRFOIL
SECTIONS
FROM TEST$ IN VARIABLE -DENSITY WIND
dC^
TABLE IL-SLOPE OF LIFT CURVE, ae= dao (PER DEG.)
Thickness
designa-
\
b
C
b
06
gg
12
15
Is
21
25
0.102
0.101
0.101
0.100
0.098
0.094
0.089
112
57
TUNNEL
TABLE IV.-MAXIMUM LIFT COEFFICIENT, CL..
Th'
d
d__
-
-
13
M
A
• 0.98
x 1.27
-1.53
1 1.53
x 1.49
12
Z
,
or des.
tgnntinn
00 .---..-.._._
011 -___________
22
--.-l0,3-- -".-1,0,2-- 7:lot
-10- 2- - ------- ------- -------:f03
:097
1
l
03
02
099
ON
om. -----02
25
26 ____________ ------- ------- -27____________ ------- ------- -23::::::::::::
42- ----------- ------- ------- ------- ------- ------- -I :::-,
43---- ------10%
: ,1 21
'm
: 0,03,
44 ------------ : 102 : 183
.101 - .096 .095
46____________ .104 .103
------- - --0
47____________ ----- -------------- ------- ----- - ------- ------82.....-_____
-------- - -1
84
----- --
.
:1
.104
H6 7 - --- ig - ------.096
.102
OD9
.099
: OD6
.101
l04
.101 .103
.101
. 099 .095
094
.103
.101
1
101
02
00
00
: 'OO'D
. 1,30̀
:017
098
.097
:10I
: 101
101
OK
1.67------------ --- - -- --- - --I
0.101
------- ------- ------- - ----- - - --- ------23 ..-______.__ • 1.04
-1.61
11.60
°1.54 ------- ------24 _____.._---- • 1.01 11.51 1 1.59 1 1.55
1.43 -1.36
25 ----------- • 1.03 -1.38 1 1.60 1 1.53 • 1.48 1 1.38
- ----------- ------- -------------- ------- -------------27------ - --- ------- ------- ------- -------- ------ -------
1
42 ____________
43 _______..._.
44
45 ............
46
47 ____________
------- -------I1 : M
160
1:
60
.•11.56
....
:11
----
------ ------- ------lf : 51
11-41
b1 : 63
^1.61
17
.• 1.47
1
. 1.69
62
A 4-..
.
.
Thickness
00 ___-__._.___
-0.1
0.0
22---- - - ---- --------23. __________-1.8 -2---. 0
24. __________- -1.7 -1.7
25_ -____.___-_ -20 -2.0
26 ----------- ------- ------27 ------------ ------- ------2
4
12
15
18
25
12
Theor.
0.0
0.0
0.0 -0.1
0.0
0.0
- 0go
9
'
-1.7 -1.7 -1.9 -1.7 ------ - 1.8
-2. 0 -2.0 -20 -1.8 ------ - 2. 1
3
- ::::: :::-:: :::::: :..
-N
-1.
-2 08
-2.29
2. 59
-3.04
21
-- - ----- - ----- ---- -----
------- ------- ------ ------ ------ ------ ------
- 3 .4
43 ------------ -3.8 -3.6 -3.7 -3.6
3
8 -3.5 -3.6 ------1
_,: 9 -3. 4 --8.1 1.
4.
-3. 9 -3.
46......... . _.
3
4.1 - 4.
-4.1
-3. 4 ...
-4.2
40_ ___________ ------- ------- -4.6
-5.0
47 ____________ ------- ------- 92
----------- ------- ------- ------ ------ ------ ----------b.2
-5.4 -5.4 -5.4 - -8.2 -6.2 ----1-7 -I 2
-5.7 -b.3
14
-1: 6 -1.1
-6.1
-^I: 7
_
_
66...._._.. _. -8.3
0
6. 3
6.
66.__________ ------- ------- -- :::- :::-::
87_ ___________ ------- ------I
Based on straight portion of
lift curve
- 6.2
-5.5
-1-7
2
6
7.0
-3.60
-3.84
.15
-4.58
-5.18
-6.09
-5.40
-5.76
-6.23
-6.88
-7.78
-9.13
extended. Bee curve for actual value.
11.60
1
2
'1:
11.02
------- ------1: 1
:l
• 1.71
11
11
1 1.69
1171
1. 82
•
1.417
----- 6
.....
------
:168
1: 8
•
097
09
-1.53
-
62_ ___________ ------- ------- ------- ------- ------- ------- ------1 : b4 11.87 1l.64 ^l.65 :,1: 43, 1 1.37 :::::::
64
-: 1 41
85
11
65 ___._ _ ___.__ -1.29
43 1 1: 6781 •1: 75 -1:89
167
-1.61
^•1.49 -------66 --- ------- ------- ------- ------- ------- ------- ----.......
.- 67 ............ ............ ...... ........ .......
Additional tests to determine variation with camber.
06
1 1.2()
-------::::-:
------- ::::::
26
TABLE III.-ANGLE OF ZERO LIFT, .4 (DEGREES)
desiret on
\
b ell
desX
ig-tion
- Laii
22 ___ _________
NOTE-I,ettff indicates type of lift
enrva,
peak.
1 1.75
11: 11
<I.
1 1.75
6
1.83
95
58
REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
TABLID V.-MOMENT COEFFICIENT AT ZERO LIFT, C.,
Thickness
06
09
12
16
-0.002
-0. ON
-0.002
21
18
25
12
Theor.
ber des
ignation
00_. ._.____.
2
2
23 -----------.__.._.____.
24 __._-______
25 -- - ---- ---28_ ___________
27 ____________
41 ___________ _
43 .___.__.____
44 -----------46 _-._.__-___
46 ____________
47 ____________
62 __ _ _________
63__ _ __ .__-__.
64 . . _ __ _.___._
65_ ______ __._
-0. 002
0.000
-
0.001
-0.002
-0.029
-.038
3
159
.158
-.154
x-.160
-.139
-.129
86_.________
67 ----A
-0.003
------------ ------------ -----------------------.036 ------------.038 -----------01 6 ------------.034 ------------ ------ ------ ------ ------.039
-.044
W
-.010
-.037
-.036 --- ---------.048
­ 062
-.050
-.019 -.047
-.043 -----------____________ ---- -------- ------------ ------------ ------------ ------------ ----------------------- ------------ ------------ -- ---------- ------------ ------------ ----------------------- ----- ------ ------------ ------------ ------------ ------------ ------------.075
-.073
-.072
-.008
-.066 -.057 -------- ----.087 - 086
-.083
-:078
- . 087
-.071 ---- : --- : --- 109
-:106
-:102
-.097
- . 094
-:082 ---- --- --____________
------------ ------------ ------------ ------------ ------------ -------- -------------- ---- -------- --------------- ---- ---- ------ ------ ------------------------------------- -------- -------------------------------III
---- -----------110
il
-:106
-:097
I
-.090 -----------29
as
29
-.125
-.118
-.110 .... -------
portion
- . 076
-.059
-.075
-.039
-:105
-.124
-.143
-.0739
-.0994
-.1062
-.1257
-.1497
-.1825
-.044
-.054
-.000
_.087 -.1109
_.II0
-.1342
-.132
-.1594
- . 169
2W
:186
-.206
2737
__
----------
Based on straight
0
-0.0370
- . 0447
-.0531
om
0749
-.0912
of moment curve attended. See curve far actual value.
TABLE VI.-DISPLACEMENT OF CONSTANT MOMENT
POSITION IN PERCENT CHORD AHEAD OF QUARTER-CHORD POINT (100 TIMES VALUES OF n FOR
EQUATION Cm 04 = C,+nCL)
TABLE VIII.-OPTIMUM LIFT COEFFICIENT, C,.,
Thk__
Td.
design s
an
th"'
Cam-
ThickThickness
Coca ion
00
09
12
IS
15
21
25
12
go
\d^
bar
am
i
if..
• designsber des-
is
gnatim
Ig
-
_____.______ 0.00 0.00 0.00 Q. 00 0.00 0 00 0 0D0.00
___________ -- ---- ------- ------ ------ ------ ------ ---- - .17
:: _ ::::::: :17
: 1171
15
24
......
. 21
:11, :1If0----- ------ ------ : 20
25--
18
.16 . ,
IS
, :11
08 :03, - ::::: .23
20 ____________ ------- ------- ------ ------ --- -- ------ ------ .20
27____________ ------- ------- ------ ------ --- -- ------ ------
.20
42 ____________ ------- ---- _ _ ---- ------ --- -- ------ ------ .36
93 ____________
.35
.30 .23 .12 .20 .05 ------ .34
: 1
:36 : 33 : 22 : 16 : 10 _ _: -- - : 1
to
34
27
22
16
30
40 _ -_________ ------- ------- ------ ------ ------- ------ ------ .30
47 ---------_ ------- ------- -- --- ------ ----------- ------ 62 ____________ -- ---- ------- ------ ------ ------ ------ ------ .55
63
63
40
M
24
13
47
6460
:
:4
6642 _^:: 33
^15
---- - - j 445
j 21
10
-------
----66____
.60
.53 .42 .33
_gg
00
fgnation
00 _--_.____-_0. 7
0.7
0.9 1.1
22
23::::::::::::
------- ------3 -- ------3 -- ------5-3
24_ ___________
.1
.3
.4
.7
25 .-_______ ._ .0
.2
.6
.3
26 ----------- ------- ------- ------- ------27 ----- ----- ------- ----- - ------- ------42 ------ ---43_ ____ _______
44_____ _______
46 :: --------
1.4
IT - a - ed 0.9
1.0
1.0
1.5
1.7
.......
-------
.0
.7
.8
.8
-------------
------- ------- ------- ------- ------
.5
1.2
.4
.7
1.6
.3
5
L4
1,7
. 10
1.0
:5
.3
:-8
:9
1.4
1.7
- ---------------- ----'
I:
------ ------.2
.3
.3
-1
46_ _____ _____
47 ___ _ __ _____
62
.
.
22_
23-
08
.33
1
.
1
64_ ___________ ­ 7
.0
.0
86___________ _
.0
66__ ___ __ _
67:-:.::.. 1^^ --------
.6
.7
.9
1.6
---------------
1.3
1.8
1.7
1.9
-------- ----
-------------
.8
1.5
1:::::::Il
2: of
II.-MINIMUM PROFILE-DRAG COEFFICIENT, Cna_
Thickness
_ des na
06
'1"
i
9
12
15
18
21
ig
25
0
12
.272
.616
.544
512
:50
.512
.51A
.923
81
:117
"M
so
:'
816
Theoretical lift coefficient at "Ideal" ougle of attack.
TABLE IX.-RATIO OF MAXIMUM LIFT COEFFICIENT
TO MINIMUM PROFILE-DRAG COEFFICIENT,
CL... 147DO.Thickness
d
gs
gg
ig
S.
is
18
bar
C des,
Ignation,
0.308
272
: 250
.251
.256
667::::
---6
I
TABLE
V
0
26
12
\^d
I
0.
00 ._.____.--__ 0..
22 -----------
0074 0.0093 0.0093 0.0108 0.0120 0.0143
^ 0083
00 -_-____...__
------- ------- ------- ------- --- --- ------- -------
23
24:.
26 ___________
:0073
0070
.0073
: 0083
0080
.0081
: 00ag
0090
.0099
: 0100 ------- - ----- -------
0099 .0112 .0127 ------
.0103
.0112 .0126
•0087
.0089
.0095
26
00"
27___ _________
42 _ ___________
0090
------- ------- -- ---- ------- ---- -- ------- -------
49 .____..__.__ .0090 .0089
.0107
.0119 .0134 -------
.0101
44. - ------- .0076 .0086
.0095
.0105
.0110 .0132 ... __
46-____.____ . 0037
.0188 -------
0003
.0103
.0113
.0125
------- ------- ----- - __- ------- ------- -------
46 ____________
47_ ___________ ....... ....... ....... ....... ....... ...... ----- _
:
62_ ___________ ------- ------- ------- ------- ------- ------- -------
63__ ------- .0092 .0101
.0108 .0120 .0130
0144 -------
64
:01 :1
:1 :0120 .0132
.0146 -------
0100
0111
65:
0093
0127
.0141
.0154
0092
.0096
.0092
.0995
_ 0098
.0104
185
172
184
164
138
115
84
184
22- ----------- ------- ------- --
2a ------------
142
182
24 _____.___
144
189
177
156
128
106
-------
26
180
_ _-14
__ 170
148
132
109
-------
20 ___ _ ________
----
----- - _ ----- - _ ----- ------- ------- ------- 27 ____________ ------- ------- ------- ------- ------- ------- ------- 184
181
100
184
187
187
42 ____________ ------- ------- - ----- ---- -- ------- ------- ------- 43 --__-____ ___
161
140
123
96
-------
44 ______-- ::::
1, 82
IN
170
150
127
104
-------
46. .______.___
132
169
164
143
123
-------
106
46 ------------ ------- -------------- ------- --------------------- 4> ____________ ------- ------- ------- ------- ------- ------- -------
186
172
179
178
178
175
82.-------- -iii __ __i6__ ­ iii­ _' iiF _ ____
63 - --ii6 ___ ii
64 .-___--_._-.
160
179
159
Isla
114
7 ---------
: ____-----
-
lag
171
65 ..__.....__.
158
132
114
1`
97
I'a
I.
.
0101
.0102
.0104
of.
-
A.
:g-
z
a.
x
d
t=i
x
O
C/)
^d
CL
o
^z
H
w [
o ^
w
m
o
D 0
y
Q
PV
cn
CD
o
,^
^o^
p ^
Z
d
>
vn
H
00
Cyi7
O
r
A
"^
,^„^
2
^'
"^^^rt'