The characteristics of 78 related airfoil sections from tests in the variable-density wind tunnel

Document

REPORT No. 460

THE CHARACTERISTICS OF

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.

PRATT, Brigadier General, United States Army.

HENRY C.

Chief, Mat€riel Division, Air Corps, Wright Field, Dayton, Ohio.

WARNER, M.S., EDWARD P.

Editor "Aviation," New York City.

ORVILLE WRIGHT, Sc.D., Dayton, Ohio.

GEORGE W. LEWIs, Director of Aeronautical Research.

JOHN F. VICTORY, Secretary.

Engineer in Charge, Langley Memorial Aeronautical Laboratory, Langley Field, Va.

HENRY J. E. REID, JOHN J. IDE, Technical Assistant in Europe, Paris, Franed.

EXECUTIVE COMMITTEE JOSEPH S. AMEs, Chairman.

DAVID W. TAYLOR, Vice Chairman.

CHARLES G. ABBOT. WILLIAM P. MACCRACKEN, Jr.

LYMAN J. BRIGGS. CHARLES F. MARVIN.

ARTHUR B. COOK. HENRY C. PRATT.

BENJAMIN D. Fo ULOIS. EDWARD P. WARNER.

ERNEST J. KING. ORVILLE WRIGHT.

CHARLES A. LINDBERGH.

JOHN F. VICTORY, Secretary.

E

REPORT No. 460

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 series of propeller sections, the airfoils have not been

systematically derived in such a way that the results

was made in the N.A.C.A. variable-density wind tunnel

could be satisfactorily correlated.

at a large value of the Reynolds Number. The tests were

The design of an efficient airplane entails the careful

made to provide data that may be directly employed for a

balancing of many conflicting requirements. This

rational choice of the most suitable airfoil section for a

statement is particularly true of the choice of the wing.

given application. The variation of the aerodynamic

Without a knowledge of the variations of the aerody-

characteristics with variations in thickness and mean-line

namic characteristics of the airfoil sections with the form were therefore systematically studied.

variations of shape that affect the weight of the struc-

ture, the designer cannot reach a satisfactory balance developed by combining certain profile thickness forms, between the many conflicting requirements.

obtained by varying the maximum thickness of a basic

The purpose of the investigation reported herein was

distribution, with certain mean lines, obtained by varying

to obtain the characteristics at a large value of the

the length and the position of the maximum mean-line

Reynolds Number of a wide variety of related airfoils.

ordinate. A number of values of these shape variables

The benefits of such a systematic investigation are

were used to derive a family of airfoils. For the purposes

evident. The results will greatly facilitate the choice

of this investigation the construction and tests were limited

of the most satisfactory airfoil for a given application

to 68 airfoils of this family. In addition to these, several

and should eliminate much routine airfoil testing.

supplementary airfoils have been, included in order to

Finally, because the results may be correlated to

study the effects of certain other changes in the form of the

indicate the trends of the aerodynamic characteristics mean line and in the thickness distribution.

with changes of shape, they may point the way to the

The results are presented in the standard ,graphic form

design of new shapes having better characteristics.

representing the airfoil characteristics for infinite aspect

Airfoil profiles may be considered as made up of cer-

ratio and for aspect ratio 6. A table is also given by

tain profile-thickness forms disposed about certain

means of which the important characteristics of all the

mean lines. The major shape variables then become

airfoils may be conveniently compared. The variation of

two, the thickness form and the mean-line form. The

the aerodynamic characteristics with changes in shape is

thickness form is of particular importance from a

shown by additional curves and tables. A comparison

structural standpoint. On the other hand, the form of

is made, where possible, with thin-airfoil theory, a

the mean line determines almost independently some summary of which is presented in an appendix.

of the most important aerodynamic properties of the INTRODUCTION airfoil section, e.g., the angle of zero lift and the The forms of the airfoil sections that are in common pitching-moment characteristics.

use today are, directly or indirectly, the result of The related airfoil profiles for this investigation were investigations made at Gottingen of a large number of derived by changing systematically these shape vari- airfoils. Previously, airfoils such as the R.A.F. 15 ables. The symmetrical profiles were defined in terms of a basic thickness variation, symmetrical airfoils of and the U.S.A. 27, developed from airfoil profiles investigated in England, were widely used. All these varying thickness being obtained by the application investigations, however, were made at low values of of factors to the basic ordinates. The cambered pro- the Reynolds Number; therefore, the airfoils developed files were then developed by combining these thickness may not be the optimum ones for full-scale application. forms with various mean lines. The mean lines were More recently a number of airfoils have been tested in obtained by varying the camber and by varying the the variable-density wind tunnel at values of the shape of the mean line to alter the position of the Reynolds Number approaching those of flight (refer- 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

the ordinates y are given by an equation of the form ordinate of the mean line as the position of the camber.

d

f y = a ^x + ax + a 2 x 2 + a 3 x' + a4x4

An airfoil, produced as described above, is designated by

a number of four digits: the first indicates the camber in

The equation was adjusted to give the desired shape

percent of the chord; the second, the position of the camber

by imposing the following conditions to determine the

in tenths of the chord from the leading edge; and the last

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 x=0.3 y=0.1 dy/dx = 0 from the leading edge, and a maximum thickness of 15 percent of the chord; the N.A.C.A. 0012 airfoil is a (2) Ordinate at trailing edge symmetrical airfoil having a maximum thickness of 12 x=1 y=0.002 percent of the chord.

In addition to the systematic series of airfoils, (3) Trailing-edge angle several supplementary airfoils have been included in

x=1 dy/dx= —0.234

order to study the effects of a few changes in the form (4) Nose shape of the mean line and in the thickness distribution.

Preliminary results which have been published in- X=0.1 y = 0.078 clude those for 12 symmetrical N.A.C.A. airfoils, the The following equation satisfying approximately the 00 series (reference 2) and other sections having differ- above-mentioned conditions represents a profile having ent nose shapes (reference 3); and those for 42 cam- a thickness of approximately 20 percent of the chord.

bered airfoils, the 43 and 63 series (reference 4), the 45 and 65 series (reference 5), the 44 and 64 series (refer- t y = 0.29690.

1 / — 0.12600x — 0.35160x 2 + 028430x3

ence 6), and the 24 series (reference 7).

— 0.10150x4 . / - ^^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 ±v = 0.29690 Arx_- 0.12600 x -0.35160x' +0.28430 x' -0.10150x' Basic ordinates of N.A.C.A. family airfoils (percent of chord) 5.0 1 Ord____.I O 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 This equation was taken to define the basic section.

tunnel of the National Advisory Committee for Aero- The basic profile and a table of ordinates are given in nautics during the period from April 1931 to February figure 1. Points obtained by removing the camber 1932.

from the Gottingen 398 and the Clark Y sections, and DESCRIPTION OF AIRFOILS applying a factor to the ordinates of the resulting thickness curves to bring them to the same maximum Well-known airfoils of a certain class including the thickness, are plotted on the above figure for com- Gottingen 398 and the Clark Y, which have proved to parison. Sections having any desired maximum thick- be efficient, are nearly alike when -their camber is ness were obtained by multiplying the basic ordinates removed (mean line straightened) and they are reduced by the proper factor; that is to the same maximum thickness. A thickness variation similar to that of these airfoils was therefore chosen for

±y,=6- (0.29690.1/x-0.12600x-0.35160x2

the development of the N.A.C.A. airfoils. An equation

t

defining the shape was used as a method of producing fair profiles.

+0,28430e-0.10150x')

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL where t is the maximum thickness. The leading-edge and radius is found to be [(1-2P)+2Px-el yr=(1

p)2

a =' .Iota r` 2 0 20 ae) (aft of maximum ordinate) When the mean lines of certain airfoils in common The method of combining the thickness forms with use were reduced to the same maximum ordinate and the mean-line forms is best described by means of the compared it was found that their shapes were quite diagram in figure 2. The line joining the extremities different. It was observed, however, that the range of the mean line is chosen as the chord. Referring to of shapes could be well covered by assuming some the diagram, the ordinate of the thickness form is yt simple shape and varying the maximum ordinate and measured along the perpendicular to the mean line its position along the chord. The mean line was, from a point on the mean line at the station along the therefore, arbitrarily defined by two parabolic equa- chord corresponding to the value of x for which yt tions of the form was'computed. The resulting upper and lower surface yo= be+ b x+ b2 X2 i points are then designated: - where the leading end of the mean line is at the origin Stations x„ and x, and the trailing end is on the x axis at x=1. The Ordinates y„ and yt values of the constants for both equations were then u l refer to upper and lower expressed in terms of the above variables; namely, where the subscripts and (1) Mean-line extremities surfaces, respectively. In addition to these symbols, the symbol 0 is employed to designate the angle be- X=0 y^=0 tween the tangent to the mean line and the x axis.

X=1 y'=0 This angle is given by (2) Maximum ordinate of mean line 6 = tan-1 dx x=p (position of maximum ordinate) y Oa ya lxa. B=Ion-' dx ./0- 9 yt v Mean /ins P1 Chord p ^ x Or /xt, yt/ x„ = x - y, sine Yu ' Y B + yt cos - ./O L Rodius fhrough end of chord xt = x + y, sin a y t = y t - yt cos e /.00 Sample calculations for derivation of N.A.C.A. 6821 z v, tan a sin v aos a y, sin e y. cos a z, x, ue v. m 0 0 ........................ 0 0 0 0 0 10.40000 0.37140 0.92840 0.03094 0.00084 0.03683 0.01438 -0.02805 0.01250 0.03314 0.00489 .38333 .35793 .93375 0.01186 .30000. .16503 .30000 -.04503 .3000 .10503 .06000 0 0 1 0 .10503 .am .07988 .04898 -.07347 -.07327 .99731 -.00585 .07906 .80585 .12863 .59415 -.0307 1 .0221 0 -.17143 -.16897 .98662 -.0037 .00218 1.00037 .0218 .99963 -.0218 Slope of radius through and of chord.

Method of calculating ordinates of N.A . C.A. cambered airfoils.

FIGURE 2. - yo=m(maximum ordinate) The following formulas for calculating the ordinates may now be derived from the diagram: dye/dx = 0

x a =x-y, sin 6

The resulting equations defining the mean line then cos 0 ya =y,+y, became x l =x+y t sin 0 cos 0 y t =yr -y,

Myr= I2Px - el

p2 Sample calculations are given in figure 2. The center for the leading-edge radius is placed on the tangent to (forward of maximum ordinate) the mean line at the leading edge.

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS A family of related airfoils was derived in the manner models, which are made of duralumin, have a chord described. Seven values of the maximum thickness, of 5 inches and a span of 30 inches. They were con- 0.06, 0.09, 0.12, 0.15, 0.18, 0.21, and 0.25; four values structed from the computed ordinates by the method of the camber, 0.00, 0.02, 0.04, and 0.06; and six values described in reference 8.

of the position of the camber, 0.2, 0.3, 0.4, 0.5, 0.6, and Routine measurements of lift, drag, and pitching 0.7 were used to derive the related sections of this moment about a point on the chord one quarter of the family. The profiles of the airfoils derived are shown chord behind its forward end were made at a Reynolds collectively in figure 3.

Number of approximately 3,000,000 (tank pressure, For the purposes of this investigation the construc- approximately 20 atmospheres). Groups of airfoils tion and tests were limited to 68 of the airfoils. Tables were first tested to study the variations with thickness, of ordinates at the standard stations are given in the each group containing airfoils of different thicknesses figures presenting the aerodynamic characteristics.

These ordinates were obtained graphically from the but having the same mean line. Finally, all airfoils computed ordinates for all but the symmetrical sec- having a thickness of 12 percent of the chord were ^----- 06 2,506 2306 2406 2606 2706 2409 2509 2609 2709 2209 2309 - G X09 2612 -^2712 '— 2212_ 2312 ^2412 001Z- X315 2615 -^ O55 8 ' ^ 2618 2918 2318 2418 0-018

C -^ ^_

C_^

' _.^1 '^ -^

L -2421

^

^ 4206 4706 4306 4406 4506 4609 4709 4209 4309 4409 4509 9212 4312 4412 4612 4712 4615 4715 4215 4315 4415 4515 4218 4318 4418 4618 4718 4221 4321 4421 4521 9721 6406 6506 6706 6206 — 6306 6606 6409 6509 6209 6309 6609 6709 6212 6412 6512 6612 6712 6218 6318 6418 6518 6618 ^-6718 6221 6321 6421 6521 6621 6721 F—E 3.—N.A.C.A. sWoll pmffiw.

tions. Two sets of trailing-edge ordinates are given. tested to study the variations with changes in the Those inclosed by parentheses, which are given to mean line.

RESULTS facilitate construction, represent ordinates to which the surfaces are faired. In the construction of the The results are presented in the standard graphic models the trailing edges were rounded off.

form (figs. 4 to 80) as coefficients corrected after the Three groups of supplementary airfoils were also method of reference 8 to give airfoil characteristics for constructed and tested. The derivation of these air- infinite aspect ratio and aspect ratio 6. Where more foils will be considered later with the discussion.

than one test has been used for the analysis, the infinite APPARATUS AND, METHODS aspect ratio characteristics from the earlier test have been indicated by additional points on the figure. Table A description of the variable-density wind tunnel and the method of testing is given in reference 8. The I gives the important characteristics of all the.,airfoils.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL T Lw0r.

/2 U i 0 0 .947 - /.307 y W - /0 /.777 y p O 20 40 60 80 100 -241 -2.673 Per cent ofch-d •l0 -2.669 2.

y 36 v .00/ .44 .09 -2.902 l a 7 -2.282 32 .40 G.OB ,832 ,y - /.3/2 - .724 u .36 u .07. 28 ^ 1.8 .403

1 0 .O

L6 .32 8.06 -0.40 V 20 w L4 .26% 0.05 28 0 .v c. p.

L D 16, /.2V .24 ^ v .04 24 0 k K 12^ 40 20 °.03 N 20^ C I E u C 8 ^^ .8 8 .16 0 .02 0 y o O l l 4,^ 60./2 .0/ o /2^ u

.4^ .0B 0 g0

-4 0 .2 .04 0 4u m v 0 0 ^-.z -8 a oQ Airfoil: N.A.C.A. 0006 R.N.:.^21Q000 .0 Sze: 5"x30" Ve%(0../sec.): 66.5 _ -4 W -/2 ' 2 . 3 Pres.(sthd.. otm.):20.8 Dofe: l-4-32 Airfoil: N.A.C.A. 0006 R.N- : i Where felted.L.MA.L. Test.• U.D.T. 744 Test: V L Date.-I-4-32 -/6 -.4 -.4 ll effect. -9 0 Corrected for tunnel-wo Corrected to infinite aspect ro 2 .4 .6 8 10 1.2 1.4 ..4 1216 20 24 28 32 0 4 8 Angle of .,tack, a (degrees) Lift coefficient, C F-- 4.-N.A.C.A. 0000.W.H.

cu sta. Op'r. C.'r.

U /0 v a O O 0 U -C O 125 1.420 -1. 961 ,96/ -1.96/ U - /0 5.0 666 -2666 4 0 7.5 3./50 -3.150 -3.512 O 20 40 60 80 10 .5/2 15 4.009 -4.009 Per cent ofchord 20 4.303 -4.303 h 4.456 -4.456 36 w 304.50/ -4.501 .44 .09 40 352 4.352

l

503. -' -3.97/ D 60 3.423 -3423 32 ag 20 .40 ^'.. 08 70 2.746 2.746 60 /.967 -/.967 S N 901. 066-1.066 U ea 1.8 .36 .g,.07 100!095) (A 5J /00 0 0 1.6 .32 0.06 L.E. Rod.: 0.69

l

e U 20 w 1.4 .28- 0.05 28 1 C 1?

c. p.

/6 p

2421 1.2 .24 v.04 G ^ v

/2

/.0 8.208 Q.03 q 2004/ LD u

.88.160 .02 8'^

o /6D 6i t ^ 80 k . .0/ 4^ 6 .12

a /2^Bi

w Op 0 •4`1 .08, 0 8 /0, 11 vi I

-4 0

.2 .04

° 4u

C .o y -.2 0 0 "a c

Airfoil: N.A.CA. 0009 R.N.:3,2/0,000 e

°u c Size: 5-x30" Ve/.(Alsec.):68.5 -4 -.2 ^ -.3 Pres.(stnd.otm.):20.8 bbafe: /-6-32 F Airfoil.• N.A. C. A. 0009 R. N.: 32/0,000 Where tesled:L.M.A.L. Test:VOT 746 Dote. /-6-32 Test: V.D.T. 746 -.4 0 -.4 Corrected to inf)iVM aspect ratio

Corrected for tunnel-wall effect. -8 u

0 .2 .4 .6 .8 10 /.2 14 12 16 20 24 28 32 -4 -6 -4 0 4 6 Lift coefficient. Cl Angle of oftack, a (degrees) Fla". 6.-N.A.C.A. 0000 MIMI.

8 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAIITICS

Sla Up'r.

./2 v o /o 0 0 0 t u O /.25 /.8.94-1.894 2.5 26/5 5.03. 555 -2615 -3555 o 7.5 4.200 -4.200 O 20 40 60 00 /0 4.683 -4.683 60

./0- /55.345 -5345 Per centofchord

20 5738 -5.738 25 5.941 -5.941 306.002 -6.002 .44 .09 40 5.603 -5603 50 5.294 -5294 604.563-4.

2.0 .40 -08 O 70 3664 563 -3.664 t 80 2623 -2623 N 901.445- 1.446 /.8 .36 x.07 'w 95 .807 - .607 %00 (. ^) (-. p 6) a 1.6 .32 0.06 L. E. Rod.: 1.58 C v C1 U 26 0 /.4 .281 0.05 ,m V e.p.

24 0 20

v.04 v ^=

-0 40

1.01.208 .03 L/

p /660

A .16 02 o O 12 80 ..6'.12 .0/ 9 -/00 .4'.08 0 i

.2 .04 ° 4 u

o u

0 0 0 Q

u .2 Airfoil: N.A. C.A. 0012 R.N:3,230,000 u c -4 4 Size: SWO" V.I.(k/sec.f: 68.4 -.3 -.2 +^ Pres. (sfnd. otm.):207 Date. , 12-30-31 Where tested. • L.M.A.L. Test. , V.D.T. 743 -.4 -.4 -8 u Corrected for funnel-wol/ effect.

-8. -4 0 4 8 /2 16 20 24 26 32 ` -4 .4 .6 8 1.0 /.2 1.4 Angle of oftoch, o: (degrees) Lift coefficient. C.

FI°URE 0. - N.A.C.A. 0012 Oit 0.

a 20 Slo Up'r. L'w'r ./2 0 0 0 N p /0 U l [252.367 -2367 0 0 2.53.268 -3268 5.0 4.443 -4.443 7.5 5.250 -5.250 00 /05853 -5653 20 40 60 80 /5 6.68/ 10 -6.681 Per cent ofchord 207./72 -7./72 25 7.42 7. - 7.427 30 50 -7.502 .44 .09 40 7.254 -7.254

U $ 50 6.618 -6.618

60 5.704 -5.704 20 .40 r 08 0 70 4.S80 -4.580 C BO 3.279 -3279

v

U 90 /.8/0 -/.8/0 /.B .36 `.07 9 /.006 -/.008 `o l00 (./SB) (-.

V 1561 / O 0 1.6 .32 0.06 L. E. ROd.: 248 C 0 u 280 0 1.4 .26- 0.05 .v c. .

24 0 20 42 .24•^ wy.04 V w g20 40 ,.0 208 .113 / L 0 /6 60 .8 0 .16 .02 w l ^ a o a /2 80 .6 u .12 .0/

v

8,0100 .4 v .08 .0

t

0 4v

.2 .04 j-./

.g u l

v

O 0 y -.2 Airfoih .A. 0015 R.N.:3,200,000 C

u

Ve/.(ft./sec.): 68.4 -4 0.

-.2 3

fm):2/.O Dote:/2-9-31 . • LMAL. Test: VD.T. 728 -8 u

-,4 0-.4

or tunnel-wall effect F -6 -4 0 4 6 12 16 20 24 28 32 4 Angle of ot/oc/r, of (degrees) Lift coefficient C Flo- 7.-N.A.C.A. 0015 etrlall.

9 DENSITY WIND TUNNEL

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE - >a Up'r. L'w' c U /O 48 /2 ^c O 10 264/ -2.841 Z53.9,11", 2 -.?.922 v ^ _/0 44.

SO 5.332 -5.332 C o 6.300 to -6.300 -7.024 0 40 60 BO /0 /0 7,024 20 6.0/8- _6.0/6 Per cent ofchord

20 66 6.606 1

-6.9/2 2569/ 30 9.003 -9.003 360) 44 .09 40 6. 041 -6.705 l 507.84/ -7.94/ G 60 6845-6.845 2.0 .40 08 32 70 5436 -5.496 C 60 3935 -3935 90Z/ 72 -2.172 28 ti /.8 .36 x.07 95 1.210 -/.2/ 0 (./0 89) ,O (-.0 9) AX z a 46 .32 0.06 24 p C

L.E,Had: 3.56 l

v 1 U 1.4 .28c 0.05 20 26 0 u c. p.

/.21.24 x.04 /6 p 24 L. 20 a M /.Oy.20^4.03 /2 ig y20x40 L/D U a 8 l .8.160 .02 o /6A 60 l k .6 u .12^ .0/ 4^ 01280 q - .4 -4 .08 0 Op 8 0100 0 4m -4 0 .2 .04 C, v .o l -8"' 0 0 -.2 0 Q Airfoil:N.A.CA.00/8 R.N.:,1150.000 Ue(.(H./sec.): 68.8 -,1'Cry Size: S"x30" Pres.(stnd. atm.): 20.9 Date: 1-6-32 Airfoil: N.A.C,A. 00/8 u Where tesfed:L.M.A.L. Test. • VD.T, 747 Dote: / - 6 - 32 -8 -.4 0-.4k Corrected for funnel-wolf effect. Corrected to infinite a4 2 .4 .6 .8 /.0 -8 -4 0 4 8 12 16 20 24 28 32 -4 AnO/e of oitock, a (degrees) Lift coeffi cient.0 Fionaa 8 .-N.A.C.A. 0018 airfofl.

L'wr.

St. up'r.

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 20. 40 b'0 80 t0 l0 6./95 -6./°5 /59.354-9.354 Per centofcho rd 20 to 1 -/0.04/ 25 397 -/0397 .44 30 5 -/0.503 40/0/5 -/0/55 - 50 9.26 -9.265 60 7.966 -7.986 R. .40 l 70 6.4/2 80410 -6.9/2 O -4590 90 2533 -2.533 u /.8 .36 95 /.4/2-/.4/2 0 O 1011(.221,1-.2211

/ 0 O a

/.6 .32 c L. n Had.: 4.65 N C 14 .28c 2B 0 0 .v c.p.

1.21.24 24 a 20 tE k v /.0y.20°u V20 40 L/D 0 u

s .l6 0

/6t 60 o O l .60 t: ./2 /2480 .4-'.08 60(00 .2 ,04 4 u a .o l 0 0 C Q R.N, 3,190.000 Airfoil: N.A.C.A. 0021 C Size: 5-"x30" t/ei.(H./sec.): 68.6 -2 Pres.(stnd.atm.):20.8 Date:1-7-32 Where tested'L.M.A.L. Test: VD.T. 748 -.4 Corrected for tunnel-wall effect. -8 u -8 -4 0 4 6 /2 16. 20 24 26 32 Angle of attack, of (degrees) Lift coefficient, FM 9.-N.A.C.A. 0021 al/3olL 27077 0-35--2

io

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Up'r. L'wi: z D /0 d V1 R 2S 5^7 -5497 5.0 7.406 - 7.406 C o-/0 7S 6.750 -6.750 /0 9.756-9.756 O 20 40 60 BO /L /5//./36-/1./36 Per centofchord 20//.953-/1.953 25/2.378 -/2.378 30 2504 -/2 504

.44

40 2.09 -209 50/1.029-/1.029 60 9.507-9.507 l 2.0 .40 70 7.633-7.633 O 80 5.465 -5465 t U 90 30/6 -30/6 1.680- /.68 1.8 .36 O 100 (.262) f-.262 0 0 0 U c 1.6 .32 LC. Rod.: 6.86 v

G°

28 1 /.4 .281

r

c.p. IC. 11 24 21 L 2,j. 24 r - ' k v q2004( 1.0 ,y .20 16 61

A.16 v./6^

0 0

0 /281 .6 u./2

B o /Of .4 08 ti 0 4 m .2 .04 C .o u l 0 0 R 4 Airfoil: MA.CA. 0025 R.N.:3,230,000 .0 Size: 5"x30 •• Ve/(fl/sec.): 68.3 4 0.

-.2

Pres. (sfnd ofm.J. • 20.9 Date: 1-8-32 V Where tested L.MA.L. Test: VD.T. 749 -8

-.4

Corrected for tunnel-wall effect 8 -4 0 4 8 12 16 20 24 28 32 An of attack, a (degrees) Lift coefficient, 1'..30.-N.A.O.A. 0026 airfoil.

5/a. Up'r. LWr.

D 20 y o !0 2 0 - 48 O V t 252.44 -/.46 0 25 3.35 -/.96 l p -/0 5.0 4.62 -235 ll 44 R 7.5 5.55-2 /O 6.27 9 -3/ O 20 40 60 60 11.

/57.25 -3.44 ./0 Per centofchord 40 20 774 -3.74 25 793 -3.94 30 7.97 -4.03 44 .09 36 7.66 -392 50 7.02 -3.56'

a

60 6.07 -3.05 2.0 .40 70 4.90 -2.43 .08 32

60 352 -1.74 g

d 90 /.93 - .97 95 1.05 - .ss /.8 .36 .07 28 O (./3) (-613) 00 D a 1.6 .32 .06 0 24 L.E. Rad: 1.56 W U 5tope of -d-0 0 1.4 .28 0 x .05 w ^ c.p

L/ 24 , b 20

2j .24 v .04 v

020040

101 20 ^ o .03 /2 l

p /6^ 60

B ` .16 0 .02 8' l l O O 0 12 80 6u./2 8'-/00 4".08 0 p k c a 0 4 v u. .2 .04 C o l _.2 0 0 w R 0 Q -8 Aif.il.-N.A.C.A.

2212 R.N:3,220,000 0 Size: 5"k30" V IMIsec.):68.4 _4 c0.

-.2 .3 P e5.(5ti7dolm.):208 Dote:12-2-31 `c Airfoil.-NA.CA. 2212 R.N:3,220,000

u

Where tested:L.MA.L. Test.-VDT 719 _4 Dofe:/2-2-3/ Test: VD.T. 7/9 -8 -. 4 -- -lB Corrected for tunnel-watt effect 0 j Corrected to infinite ospecfrot/o -8 -4 0 4 6 /2 16 20 24 28 32 .4 -2 0 .2 .4 .6 .8 /.0 /.2 /.4 t.6 18 Angle of ottack, a (degrees) Lift coeff/cient,C FIGME 11.-N.A.C.A.

2212 airfoll.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE - DENSITY WIND TUNNEL 11 5^ up'r. L'^ r.

D /0 2 48 Z25 1.16 - .73 ° g 0 25 /.70 -.95 ly / 44 ^ 0-/0 5.0 2.43 -/./5 7.5 30/ -/.22 1 0 20 40 60 80 /0 /0 3.46 -1.22 /0 40 /5 4.18 -/./6 Per centofchord 20 4.65 -1.09 25 4,91 -1.04 .09 360 30 S.00 -1.00 .44 404.86-.94 l 50 4.49 - .8/ V s0 3.92 - .65 .06 32 D 2.0 .40 70 3.19 - .48 60 230 - .33 v U 90 /.26-.19 .07--28 b /.8 .36 .9 95 66 - .13 l00 (.06) (-.06) O /00 - 0 .O6 24 0 /.6 .32 0 c C3 C.E. Rod.:0.40 Slope afrod'us lhraugh end of .OS 00, 1.4 .28 P 28 0 chord: 2115 N 0 N O .04 /, 2 24 20 /6 0 ti .24 C. .

a m '^ 03 z /O.w.20$ A20`40 LD C w

.02-a 118 eo./sa

g /6 60

u .0/< .6^./2 u /2 60 v u U .

0 0^ .4 ^ .08 8'-/0 O Y -4 0 .2 .04 V` ° 4 v Ca v v 0 0 v Airfoil: N.A.C.A. 2306 R.N.:3,080,000 .c Size; 5"X30" Vel(ft./ser,4:69.8 -.2 .3 -4 0.

Pres.(sfnd.otmJ206 Dofe:3-24-31 c Airfoi/. N.A.CA. 2306 R.N.:3080,'00 -/2 -8 u Where 1-1.1. L,M.A,I, Test: l<D.T. 660 _ q 9ote:9-24-31 Test:. V. D. T. 680 _/6 -.4 Corrected to infinite aspect ^-otio Corrected for tunnel-wall effect

-4

F -4 .2 0 .2 .4 .6 .8 10 /.2 1,4 /.6 18 -6 -4 0 4 6 12 16 20 24 28 32 Lift coefficient.0 An of attack, a (degrees) FIGURE 12 .-N.A.C.A. 23N airfoil.

St. up'r, L' ^r.

m^ l0 I25 1.69 -1.16 U t 2.5 2.39 -/.56 y _/0 5.0 336 -2.0/ C 7.5 4. 9 -224 7 0 20 40 60. 80 /0 to a. 7 -2.36 15 5.54 -250 Per centofchord 20 6.06 -2.52 25 6.37 -251 30 6.50 -250 .44 40 632 -2.39 50 5.62 -2.13 60 5.07 -1.76 R. .40 70 4./1 -1.36 O 80 2.96 - ,97 90 /.64 - .54 U /.8 .36 95 .66 - .33 0 O 100 (./01 /-./O) loo - 0

a

1.6 .32 c L.E.ROd.:0.69 C W V oJrod'us 5""' Through end of /.4 .261 280 0 chord: 2175 .v 24 0 20 c. .

w 1.0 y.20 '20 Q 40 L/D u ./6o $1660 8w 0 0

.6,"./2 F12880

.4 ^ .08 Bolo 0 o qm .2 .04 o U 0 0 0^ 2309 R.N.:2,970,000 Airfoil: N,A.OA.

.0 Size: 5'x30" l/e%(R/sec.J: 71,0 -q 0.

-.2 Ires.(st7d atm.J:203 Dote:9-24-3/ Where tested.•L.M.A.L. Test. l!O.T.68/ -.4 -8 Corrected for tunne%wa//effect -8 -4 0 4 8 /216 20 24 28 32 Angle of attack, a (degrees) Pim. 13,-N.A.C.A. 2909 airfoil.

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS St.. up , ,. L'w'r.

./2

wo 0 - 0 /.25 2.24 -1.57 2.5 3.//

m k _/^ o• x Test A -2/6

50 4.3/ -285 y 7.5 5./B -326 10 5.86 0 20 -352 40 60 80 /L 7 /5 6.89 -3.82 Per cent ofchord 20 ].54 -3.94 25 7.88 -3.99 30 B.00 .44 .09 36 v 40 ].77 -3.64 50 7.14 -3.45 l 60 6.21 -z.92 2.0 .40 .V^-.OB

0 70 5.02 -231 32

80 3.62 -/.63 90 200 - .9/ U 95 1.09 5Z /.8 .36 ^.07 28 i O /00 / i3 / 0 3 C L6 .32 0.06 L L. Rod.: 1.58 Slope ofrod'us l VQ U r of 28 0 1.4 .26" : ^.05 /harddh2%/d 20

.d

24 0 20

l2V .24 v.04

/6 p

LID k Q 20 Q 40 40 .20 1)^.03 C ° p 16, 60 0 .8 .16 .02 w l2 80 g .6^ .12 .0/

u

6,0/00

.4' 4 .08 0 Ox

°

.2 .04 G

4u 4 0

w k

0 Q

O 0 -.2

Airfoil: N.A.CA.2312 R.N.:,1120,000 u

C e

-4 0. S"x30"

Size: Ve%(N./sec.): 69.2 Pres.(sf'nd.otm.):20.9 ofe:12-2-31 Airfoil: N.AC, 23/2 R. A.N.: Where tested.' L.MA.L. Test.• V.D.T. 720 Dote: 12-2 3l -B Test:

-.4 Q -.4

Corrected for tunnet-wall effect Corrected to infinite aspect ' -8 -4 4 8 /2 /6 20 24 28 32 -4 O .2 .4 .6 .8 1.0 /.2 Angle of attack, a (degrees) Lift co efficient, F-B 19.-N.A.C.A. 2312.W.R.

Sla. Up'r. Lw^.

o - 6 - /o v o 2 U ,^.

/.25 200 -1.,96 0 2.5 3B5 -2.74 l U 5 0 5.26 -3.66 7.5 6.28 -4.25 / 0 7.06 -4.66 O 40 60 60 10 n /5 8.25 -5.13 P r cent ofchord e 20 8.97 -538 25 9.3 -5.48 30 9.50 -5.50 44 .09 40 9.22 -5.29 36 v 50 7.47 -4.77 O l 60 7.36 -4.06 G 2.0 .40 70 5.95 -3.22 .OB 32, 60 4.29 -228 LL EE C 90 1.36 -/.26 95 L30 .72 ( /.8 .36u.07 O /00 /.161 (-./6)

t

i6

/00 - O

U all

1.6 .32 p,06

L.E. Rod.: 2.47 C

v

5/ape afrodiu5 V u 2B 0 1.4 .28-^ p.OS cho(dh^sot .v. a

0 2 24 0

42V.24 •^ v.04 c.

/6 0

k w

v

' 20o4 0 /.0U .20 u ".03

/2 c c l L/D p D /6 6 O .8./60 .02 8e 0 0 D o ° 6 /2 .6 u ./2 .0/ u

8 /00

.4^ .08 0 0 4v .2 .04 Ok -. / o C 40 u k W O

0 4 0

w -.2 •8 ^` Airfoil: N.A.C.A. 23/5 R.N.:3,06C108Ci 0 e U Size: 5"x30" Ve1.(fl,/sec.): 69.8

-4

Pres.(s%'d.oim): 2Q8 Date: 9-25-31 Airfoil: N.A.C.A. 2315 V Where tested.• L.M.A.L. Test: Vb. T. 683 _ _ Dote: 3-25-31 -8

Corrected for tunne/-wo// effect .4 .4

'gCorrected to infinite os

-8 -4 0 4 6 /2 /6 20 24 28 32 .4 :2 0 .2 .4 .6 .8 40 Angle of allock. a (degrees) 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. 10 -B, - 0 t 0 1.25 /.// -0.60 Z. /.57 -/.04 4a- /0 5.0 2.28 -1.29 p 00 324 - 0 20 40 60 80 / 10 45 15 3.90 -1.44 Per cent ofchord 20 4.37 -1.37 4.69 -1.25 N 25 01 30 4.86 - /.12 44 .09 404.90 - .90 4.60 - .70 I f U 604.08 - .49 2.0 .40 06 70 3.33 ° 80 2.44 - 20 v 90 1.35 - .1 / U x.07 /.8 .36 1061 0 (-.061 l.6 .32 0.06 0.40 N C1 U 51ope ofnvdi,5 through end of 1.4 .26c th 20 v 2B p 0 chord: 2120 I I 1.2,j.24.04 /6 p 24 20 v /.O,n .2011 1 y -0Q40 2.03 /2 LID C u .8 .16 .02 /6D 60 m 0 0 v .6.".12 .0/ Q12-c o 80 4u .4'.08 0 8'.100 s ^. c .2 .04 C-./ -4 0 ° 48, v O 0 y -.2 -8 0Q Airfoil: N.A.CA. 2406 R.N.:3, 120,000 C u Size: 5"x30" 3 e/.(ft./sec.): 69.3 -4 0.

.2 _3r

Pres. (sfnd. otm):20.7 Dote: 9-8 - 3/ Where tesled. • L.M.A.L. Test. V.D.T. 665

,.rreC.A. -.4-.4

'B U Corrected for tunnel-woll effect -8 -4 0 4 8 12 16 20 24 28 32 c4 c2 0 .2 .4 .6 .8 /.0 /.2 /.4

Angle of ottock, a (degrees)

Lift coefficient.

FIGURE 16.-N.A.C.A. 2908 MEN].

5^. Up'r.

LW -,'b l0 1.25 /.62 U g O -123 2.5 227 -166 5.0 3.20 -2./5 Y v -/0 3.87 7.5' -2.44 /O 4.43 -2.60 O 20 40 60 60 /G 15 525 -2.77 40 Per cent ofchord RC 5.81 -2.79 25 6./8 -2.74 30 40 6.38 -2.62 44 .09 US S. -2.35 50 605. -2.02 22 . -/.63 2.0 .40 x.08 O 704 27 -124 C BO 3.10 - .85 i 90 1.72 - ,47 /8 .36 x.07 95 .94 - ,2B O /00 (./0) (-./0) too - o

L6 .320.06 c

G.E. Rod: 0.89 ^q U 0 C / Ih10 3 gh end of 28 p 0 1,4 .28 0.05 chord: 2/20 N ^ 24 0 20 L2,j.24M \.04 166 c. .

k v 0 20 40 L0.20 ,03-- /2 c 0 L w .8 W .16 /6a 60 .02 0 0 /2g 80 .6" .12 .0/ 4s U 8'.100 .4".08 0 '• C .2 .04 OE-./ -4 0 ° 4 u a .6 P 0 m -. 2 Airfoil., NA.C.A. 2409 R.N.:3,1 /0,000 Q C °u Size. 5"x30" Ve).(ft./sec): 69.3 -4'4 Pres.(sfnd.otm.):20.8 Dofe.3-9-31 110,000 v U Where tesfed:L.MA.L. Test.-VDT. 666 9-3/ D.T. 666 -B -.4 0 -.4 Corrected for tunnel-wall effect -6 -4 0 4 8 /2 /6 20 24 28 32 -4 :2 0 .4 .6 .8 /.0 Angle of ottock, a (degrees) Lift coeff/cient,C FIGURE 17, -N.A.C.A. 2109 Wrtoil.

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS So. w, V s° 0 !25 2.15 -465 2.5 299 4.13 -227 vk-/O 50 -3.01 p 7.5 4.96 -346 / 10 5.63 -3.75 0 20 40 60 80 /6 15 6.61 -4.10 Per cent ofchord 20 7.26 -423 25 7.67 -4.22 30 7.68 -4./2 .44 .09 36 ^ 40 7.60. -3.60 50 -3.34

a U

60 6.36 -2.76 2.0 .40 .08 32 70 5.18 -2.14 -/.50 60 3.75 d 90 2.06 - .82 48 .36 x.07 26 2f 95 1.14 - .46 gi - 0 /.6 .32 o.06 ' L.E.Rad.: 450 24 0

v l

Slope ofradius V U Yh -d.- end of 28 p 0 1.4 .28- 0.05 ?0 w chord: 2/20 •- 24 20 /.2V.24^ v.04 16

8 -

a k

v^

/.Ov.20 1 2.03 /2c V201 40 k u 0 /6t 60 .8 u .16 .02 L

l l o O

k o /2.2 80 .60 .12 .0/ 4 ^.

8 0 /00 .4 v .08 0 OL ° k .2 .04 V`-.l -4 0 ° 4u C .o k 0 0 -.2 v C 04 Airfoi/: N.A.C.A 2412 R.N.:.250,000 e .c U Size: 5"x30" Ve/. (ft./sec.): 66.0 _4 0.

.2 •3 Pres.(stnd.otm.): 2/.0 Dote:/2-3-31 Airfoil:N.A.C.A Where tested.• &A.L. Test: V.O.T. 721 Date: 12-3-31 Test: V. 721 16 -4 -6 U 0-4 Corrected for funnel-wall effect Corrected to . sct ratio

je

-8 -4 0 4 8 12 16 20 24 28 32 -4 -2 0 . .2 .4 `6 10 /.2 1.4 16 /.8 Angle of ottock. a (degrees) Lift cot ^t. C F-E 16.-N.A . C.A. 2412 aic[oll.

0 o. Up'n GOY: ^^ ^0 S7 u 0 Z25 2.71 -206 t 25 3.7/ -286 y k _/0 44. ./I 5.0 5.07' -3.84 p 7.5 6.06 -4.47 O 20 40 60 80 /L 10 6.83 -490 15 7.97 -542 Per centofchord 20 8.70 -5.66 25 9.17 -5.70 N 30 9.38 -5.62 36 N 44 .09 40 9.2S -5.25 l 50 6.57 -4.67 G 60 7.50 -3.90 32 2.0 .40 .08 70 6.10 -305

80 4.41 -2./5 5

v 90 2.45 - 417 28 tl 48 .36 x.07

I

95 1.34 - .66 100 (.16) (-.161 too - o 1.6 .32, o.06 C L.E.Rod.. 2.48 ^q V v 5/ope ofradius 1.4 .28-, 0.05 28 0 l h rdh2%20 f p 24 0 20 /6 0 c. .

k v C 20' 40 /2 QUO)

LD

g /s .8 .16 0 .OR--

v F 60 o a o /2 8 0 4^ .6 0 .12 .0/ k u 80/0 0 0^ .4 ^ .08 0

°

0 4y -4o .2 .04 j-/ o u m 0 0 v -.2 R 04 Airfoil: NA.CA. 2415 R.N.:3,060,000 e c °u Size: 5"x30" Ve/(fL/sec.):698 _4 0.

-.2 -.3 Pres. (stud. otm.):20.8 Dote:9-10-31 c Where tested'L.M.A.L. Test: VD.T.668 9-10-31 -.4 0-4-- - 8 u Corrected for tunnel-wolleffect. cted to 4 .6 -4 -2 0 -6 -4 0 4 6 12 16 20 24 28 32 Lift co Angle of ottock, a (degrees) Fmv .18.-N.A . C.A. 2415 airfoil.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL Si.. GWr up'r.

/0 L25 3.28 -2.45 u r 0 2.5 l U 4.45 -3.44 0-/O 44 50 6.03 -4.68 7.5 7./7 -5.48 7 0 /0 8.05 -6.03 20 40 60 80 /L /5 9.34 -6.74 40 Per centofchord 20/0./5 -7.09 2510.65 -7/8 3010.68 -7./2 .44 .09 36 m 40/0.71 -6.7/ l 50 9.69 -599 D 60 8.65 -5.04 2.0 .40 x'•.08 32 a 70 7.02 -3.97 O BO SOB -280 W 90 28/ -L53 /.8 .36 x.07 28 tf 95 /.55 - .67 4'.

/00 /./9) (-./9) .o /00 - O L6 .32 0.06 24 0 L. E. Rod: 356 v l of rad'us G U 5"p, m ' end a t , 28 0 1.4 0.05 20 u chord: 2/20 .28`c v 24 0 20 /.2V .24k m.04 /6 0 a.P.

w^ k v 20 40 40v.208 L..03 /z c.

N ^ u k L/D 0 6D60 .8v.16^ 1 .02-- 8.

0 0 - .6./2 .0/ a 1280 4s k u .4 08 0 8'.100 0^ `o .2 .04 Of-./ '4 0 0 4 n v v _ 8 rn 0 0 -.2 0 C w 0 Airfoil. NA.C.A. 2418 R.N.:3,060,000 c Size: 5'x30" Ve/(f1/sec): 69.8 -.2 -.3 -4 0.

Pres. (sfnd .1-):20.8 Date: 9-11-3/ %C KA.0 A. 2418 u Where tested . -L.MA.L Test: VDT. 669 9-11-31 V -.4 0 -.4 -8 Corrected for tunnel-wall effect.

clad to inhi,it .4 .6 .8 -8 -4 0 4 8 12 /6 20 24 28 32 -4 Angle of .Hock. a (degrees) Lift coeff/cie/ C n-RE 20.-N.A.C.A. 2418 airfoil.

sto. UP r. L'wr.

U c y o /0 - 0 U r 3.87 0 /.25 -2.82 2.5 S2/ -4.02 5.0 7.00 -SS/ 4 a -/0 7.5 8.29 -6.48 /0 9.28 -7./8 0 20 40 60 80 /0 151 -6 05 Per cent ofchord -6,52 20 //.59 25 3012 /2/5 -8.67 .36 -8.62 v .44 50 -7.31 U Q.

60 9.79 19 -6.17 l 2.0 .40 70 -0.67 a 80 74 1304 t 90 3.18 -1.66 U LB .36 as /00(. /.76 -1,06 ri 22) (-.22) /00 - 0 a /.6 .32 L.E.Rad, 4.85 0 v

l

Ci 5/ape trod,, a thro-1gh end of /.4 .28 28 ° 0 chord: 2120 v de h c.P. 24 , L 2V.24 0 20 0 v q20 LO,v .200 40 0 /D 0 v U .c o /sa so .6x.160 0 0 k o 12 y u ° 80 .6' - ./2 8,0100 .4.08 .2 .04 0 ° 4 u v 0 0 C 0 4 Airfoil.-NA.CA.242/ R.N.:3,000000 c Size. 5"x30" VeL (ft./sec.): 70.5 -4 0.

-.2 Pres. (sthd.Ol 7.) 0.6 Oate:9-//-3/ ti Where tested. 4.MA.L. Test: VDT. 670 -8 -.4 Corrected for tunnel-wa/l effect.

-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 t u 0 .36

a

.32 c v G 28 0 .28-, v 24 v .24.0

h

k

q 20F v.20^

.0 b, v./6o 0 /6

U.12 o /2

v 80 08

.04

.o

OQ

C

-4 0.

u -6

An of allock, a (degrees)

FIGURE 22.-N.A.C.A. 2608 airfoil.

r. L'w'r. 5/0. Up' /O

•/2

0 N p 0 D 1.57-1.27 ^$ 25 2.5 Z2 224 5.0 -2.56 7.5 3-'0 3.7 5 O 20 40 60 80 /00 B -2.76 . /D /0 W. 2 2.9e Per - a ofrhord /5 SOS __ 20 5.60-3. 02 -2.97 H 25 5.96 30 5/B -2.84 36 44 .09 40 6.27 -2A4 50 597 -/.97 60 5.35 -/.50

2.0 .40 t'..OB 32 °

70 4.44 -/.06 C 80 326 - .67 90 .36

/.B .36 3.07 28

95 .99 - .22 O 100 (.10) (-./O) /00 - o 24: /.6 .32 x.06 p G. E. Rod.: 0.69

v

S/ope ofredius

throh ug e ndof

/.4 .28 .05 got

dfi 0 chord: 2/25 G$.

/60

24 20 /.2V.24u v.04

Q

/.Ov.20, x.03 /2

Q 20040

u ^ Y.

8'^

.8,) ./6 0 .02

0 /6^- 60

.0/ 4U

. 2 l2 80 .6 U .12^

9 o /00 .4 v .08 0j

°

V r -. / -4 0

.2 .04 C"°/'

° 4u

v

0 Q

Airfoil:N.A.C.A.2509 R.N.:,R060,000 e

.c 3

_3

Size: S"x30" 3 e%(ft./sec.J: 69.8

.2 -/2

- 4

4 Tres. (sfnd. otm.J: 20.7 Dote: 30" 9-1 Airfoil:NA.CA. 2509 R.N.3.060,000

u teSMCd L.M.A. L. Test: U.D. T. 673

-.4 _16 Dote: 9-/531 Test: l!O.T. 673

E _ 4

-B

Corrected to infinite aspect ratio

Corrected for funnel-wa//effect '

-6 -4 0 4 8 12 16 20 24 26 32 c4 c2 0 .2 .4 .6 .8 /.0 /.2

Lift coeff/cient,C

Angle of oHOCk, a (degrees)

FIGURE 23.-N.A.C.A. 2600 airfoil.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE - DENSITY WIND TUNNEL 17

-/0 44 3.59 i o / / o v q 674 ^ll.1LIJ_LJJ_J-L11J .

0 20 40 60 80 /00 ssz 430 40 Per cent ofchorr( . l0 4.43

y 4.44

36 v .44 .03 %so 3.29 U oa zs3 2.0 .40 • .OB 32 $i /.97

r

1.8 .36 6.07 28 2( :42 k p /31 C a 24 0 1.6 .32 8 .06 e A a, l Ci fins of 26p 0 1,4 .26^ 0.05

24 R20 160

1.2V.241 m•04 LID

k

0 v

q20a 40 LO x.200 Q.0.3

12c a, a u l6 60

.0.160 .02

Q) 8l

I

O o o 0 /2 86 .6X ./2 .01

u

8 0 /OO

.4'.08 0 00 ^- c -4

.2 .04 0 0^H. D

° 48 v U

8 e

O p 0 0 .2

Airfoil: N.A.C.A.25 /2 R.N.:3,080,000 O C _ _ _/2 VeG/flfsec.:69.4 Size: 5"x30" -4 0.

' 2 Pres.(sfnd.ofm):21.0 Dote: /2-3-3/ v "? Airfoil: N.A.CA. 2512 R.N:3080,000 Where tested: L.htA.L. Te5t. , VD.T. 722 Date: /2-3-3/ Test: V. D. T. 722 -16

-8 u

-• 4 0 -•4 Corrected for funnel-woll effect.

Corrected /o infinite aspect ratio -.4 20 24 28 32 ..2 4 8 12 16 0 .2 4 .6 8 10 /.2 1.4 /.6 l.8 ng/e of oHock, a (degrees) Liff coefficient, C

24. - N.A.C.A. 2612 airfoil. FrovaE

u 51a. Op".

0 - U /.25 -2.11 /0 2.63 2.5 3.63 -2,94 _10 5.0 4.96 -3.96 C O 7.5 -4.60 5.9/ 80 /L /0 6.66 -5.06 0 20 40 60 15 7.75 -5.63 Per cent ofchord 20 8.48 -5.87

25 6.92 -5.92 NI#R

30 9/9 -5.84 .44 40 9./6 -535 50 6.62 -462 60 7.64 -3.76 l 2.0 .40 70 -2.88 6.28 O 80 4.57 -/.9B U 90 2.56 -/.07 .36 L0 95 L41 -.62 /oo !./sl !-.lsl /00 - 0 O 46 .32 C L.E. RW..: 2.41 N UQ S/ape of, ai- lhro ugh end of 1.4 .28 26 0 C& 2/25 N L 2V .24 24 2( c. p.

m

LO ,y .20 0 2004(

q

l l A 6D& o i ° o

.6X.12 0/28(

LD k .4'.08 k 8 010( C

0 4v

.2 .04 C .o u l 0 0 0 y Airfoi/.• N.A.CA.25/5 R.N:3,o60,000 C Size: 5"x30" Vet. fft/sec.): 69 8 -.2 -4 0.

Pres.(stndot. ):20.6 Doi-S-18-31 ti Where tested.-L.M.A.L. Test.-VDT 675

-.4 -B

Corrected for tunnel-wall effect

-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 18 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS J/ .44 .09 S

U

32 aaP .40 .08

r

v u

28 2f .36 x.07

24 p .32 0.06 v V u 28 20 m .26`c 0.05 •v a 24 lu

16 Ci.24v N.04

1-11.

k v A 20o /2 v .20 $ 4.03 u p ao 0 /6^ /6 . 0.02 m O .

o k . 2 /2 u u ./2 4u .0/ k U Bo 00 1 .08 0 -4 0 .04 0 4u .o m 0 y 0 -.2 -6", m u e c u -4 0.

-/2 Airfoil. N.A.C.A. 2518 U m .3 -4 Dote: 9-19 3/ -B -16 o _ Corrected to infinite - 0 .2 .4 .6 .8 /.0 __ .4 c2 Angle of attack, a (degrees) Litt coefficient, FIGURE 2B .-N.A.C.A. 2618 airfoil.

m o /0 u ^ 0 0-/0 0 20 40 60 BO /00 Per centafchord ^ ./G .44 .09 36 v 112` L 7 // S V s 2.0 .40 -.OB 2 B 32 O 7 5 U LB .36 x.07 6J .6

a

1.6 .32 x.06 24 0 [. Rpd.: 4.es v JPe Ofrod' 3 S ^j l 1.4 .28- 0.05 f C o^d^'z% s 20", v va

24 0

/.2V.24 y.04 168-

C.P.

k _ k

4.03 /2

U .lso .02 Bi

.e^

i

0 0 k

.6'./2 .0/

4^ 60.

.4 v .08 0

0 4w

.2 .04 OE -. /

-4 0 u CD v x l

0 0 -.2

Airfoil. N.A.C.A.252/ R.N.:,^ /30,000 e r Size: 5"x30" 3 el(ft/sec): 68.9 _ ' 2 .3 -4 Pres(sfnd. ahn.): 2/.0 Date.9-2/-31 0.

v tested.• LM.A.L. Where Test: l!O.T. 677

10ote: 9-2/ 3/

-8

-.4 -.4

Corrected for tunnel-wa/l effect

I Corrected to infinite ratio

^

-4 0 4 6 /2 /6 20 24 28 32 0 .2 .4 .6 .8 1 :4 Angle of attack, a (degrees) Lift coefficient 1'..27. - N.A.C.A. 2521 WdOU.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE -DENSITY WIND TUNNEL 19 Slo.

u U /0 0 - 0 [25 2.05 -1.73 Yo' 0 2.5 2.66 -2.37 1 U - /0 5.0 3.95 -a/7 p 7.5 172 -a66 7 10 5.34 -403 0 20 40 60 80 /C /5 6.25 -4.45 Per cenfofchord 20 667 -4.6/ 25 7.26 -462 30 7.51 -4.52 .44 40 7.56 -4.03 50 -3.36

7.23 a

60 6.56 -2.56 l 2.0 .40 70 5.56 -/.77 / 0 80 4.15 u 90 2.34 -.56 U k /.8 .36 95 1.26 - .33 O /00 Ly /.6 .32 L.E. Rod.: /.56 v V -,of /.4 .281 28 c'.9 choror 2/30 v 1.2,j.24 ;O^ c. p. 24 1 - 2( LID k v LO x.208 q20o 4( u o.

0 /sD s(

sv.ls^

0 0 o 12u 8( .6X./2 ti.

8 104 .4 08 4 C .2 .04 0 4 O 0 h 04 Airf.A.. , N.A.C.A.26t2 R.N.:3190,000 C Size: 5"x30" Ve4(ltlsec.):68.4 -q ti -.2 Pre s. (sthdotm):2/.0 V Where tested: L.M. A.L. Test: VDT. 723 -.4 -8 Corrected for tunnel-wo/l effect -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.

P /0 v L25 204 -17S U i O L5 262 -240 6. 0 3.90 -323 h 0-10 7.5 4.66 -3.76 7 0 40 60 60 /C 10 5.26 -4.12 20 /5 6.13 -456 Per cent of 'd 20 6.73 -4.75 25 7./2 -4.77 30 736 -4.67 .44 40 7.43 -4.18 7.13 -3.47 D 60 6.52 -2.6/ l 2.0 .40 0 70 5.57 -1.67 804.42 - .83 l U 90 257 - .33 1.8 .36 95 1.44 - .19 O (113) (-. 3) O 1.6 .32 L.ERod: /.58. 1 v Ci l 1h ough end.( 26 G 1.4 .28c chord: 2/35 o .W 2402( 1.2, .24 c. .

LID v /.o'.20 y 1 0 Q 41 0 /6 6C .8 .16 m 0 0 8( .6 ./2 60/a .4^ .08 0 4m .2 .04 .o u 0 0 0 p.

Airfoil: N.A.C.A. 2712 R.M-$060.000 c -4 0. Size: S"x30" Vel.(ft./sec.): 69.5 -2 P es.(sfnd. 0 Dote:12-4-31 ti Where tested.•L.M.A.L. Test: V.D.T. 724 -.4 -8 Corrected for tunnel-wall effect -tl 4 0 4 8 M /6 20 24 28 32 Angle of attack, a (degrees) Lift coefficient.( M.. 29. - N.A.C.A. 2712 sftM.

20 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Slo. L'wi-. cu LPL, O - 0 v o /0 /.25 3.04 -/.07 U t 0 2.5 4.13 -441 l U p -/0 5.0 575 -L69 h 7.5 6.% -/.66 to Z90 -46/ 0 20 40 60 80 /6 15 9.15 -1.55 Per cent ofchord 20 9.74 -1.74 25 9.92 -496 30 9.94 -2.06 .44 40 9.56 -2.05 50 875 -1.85 D -/.55

t

2.0 .40 60 70 Z 6.13 -1.21 4.39 - .BS 90 2.4/ - .50 U 95 /.3/ - .3/ 48 .36 k Li 1.131 (-63) /00 1.6 .32 L.E. Rod.: 756 v 5lope ofrodius Ci i throu gh end of

/.4 .,?8< 28 0

d: chor -4//0 ,N

2,j. 24 i? 24 20

c. p.

L/D a

k

/0^.20u 20x40

I II

A

U l .8m./60 16 60

.61" ./2 80

U .4 '.08 B 0/00 M 0 4 .2 .04 0 o u 0 O

2 0 4

Airfoil: N.A.CA.4212 R.N:3240,000 Siie: 5"x30" Vel.(f//sec.): 68./ -.2 -4. 0.

Pres.(sYnd.afm): 20.9 Dote:l2-/0-3/ u Where tel...^L.M.A.L. Tezt:VD.T. 729

-.4 -B

Corrected for tunnel-wo// effect -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.

A 20 Sta. Vp'r. L'w'r 0 12 0 - N p 48 /.25 /,39 -.57 0 2.5 2.05 -.64 ti -/0 50 3./O -.54 p

ffffH

7.5 393 -.35 /0 4.63 -.12 0 20 40 60 80 /6 /5 S. .32 Per cent of chord ./0 20 6.44 .68 25 6.65 B9 30 7.00 %00 44 09 40 6.62 /.02 50 6.33 /.03 D 60 5.56 99 2.0 .40 V 08 n, 70 4.54 .86 0 32D c BO 329 .65

w

90 /. ]9 U L8 .36 s5 .97 /5 .07 28 t 100 LO61 (-.O6) /00 O D 1.6 .32 0 06 0 L.E.Rod.:040 14:;z u W U° l 51- through end of

1.4 .28, o OS 20 m

chord: 4115

'80

L/D C I I I 0 24 , 0 2 /. 2V.24^ m c.p. .04-/6 p ti v q20 40 1.0,t .20 4 .03 0 /2'c

k

8. C o /6 60 .02 n l P O ^ o 128 d

.6x .12 O/

4 t

k

u

80/00 .4 .08 0 ------ .2 .04 C 4 U 4 0 n w OQ 0 0 v Airfoil. N.A.C.A.4306 R.N-W,90,000 c Q

°u

Size: 5'X30" VeL (f1./sec.): 70./ -4 0.

.2 . 3 Pres.(s £nd. of 3 Od1.:4-11-31 c Airfoil N.A.C.A. 4306 R.N.:3060.000 _/2 -8 U Where lested.• G.M.A.L. Test: V.O.T. 561 4-11-31 F Test: V. D. T. 561 _/6 _•4 -.4 0 Corrected for funnel-wall effect ' 4 Corrected to infhde aspect ratio -8 -4 0 4 8 12 /6 20 24 28 32 S` -.4 -2 0 .2 .4 .6 .6 /.0 /.2 /.4 /.6 1.8 Angle of attack, a (degrees) Lift coefficienl,C Fluvaa 31. - N.A.C.A. UN aidoll.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 21 cows ./2 o i0 - 34 o -l0 - %J7 44 -1.28 0 20 40 60 60 /00 - 1.0/ Per cenfofchard .l0 - 60 oi -_ so 44 09 36 p $ 29 - D - 1s 2.0 40 1.08 32 Co,

- Os

O ao w ).8 u.07 .36 28 .05 - `o w (-.09) O U 1.6 .32 24 p c 0.06 :T8_9 C cj U t rodi 3 s

28 0 /.4 .28 c 0.05 20 u f

s CL 24 0 42 V .24 v.04 16p Q) 0 0

0200 4O v.90^ .03

C

t

u p /6U i n .8 .16^ 02 8t ^ t

.6 u ./2^ .0/

4y

p /2 U k U 01) .4 .08 0

w

o qv u

J-. .2 .04 -4 0 c ^°

Ci - /

o n W O

0 2 - B

a tz 00.

Airfoil: N.A.CA. 4309 R.N.:3,060,000

C

u Size: 5'X30" Ve%(H./sec.): 69.8 -4 0.

-2 -/2 Pres(stnd. otm.): 20.B D&e.4-/3-31 ' 2 ^ ' 3 Airfoi/:N.A.CA. 4309 R.N.:3060,000 Where tested. • L.M. A.L. Test. • VD..T 563 Dote: 4-133/ Test: V. D. T. 563 ff_4 -8 -4 _/6 Corrected for tunnel-wall effect Corrected to infinite aspect ra{io ..4 0 4 8 /2 /6 20 24 26 32 -2 0 .2 .4 .6 .8 1.0 1.2 /.4 1.6 18

Angle of o{tock, a (degrees)

Lift coefficient.

32.-N.A.C.A. 4308 Wdoll. Fioun 51a. Up'r. L'wY.

U c /0 O - 0 N p L25 2.64 -1.29 U ^ 0 2.5 0 _/o 3.63 -/. 75 -2.19 5.0 7.5 S 6.22 -2.34 /0 0 20 40 60 80 /L ) 7/2 -2.39 /5 846 -231 Per cent fchord 20 9.34 -2.17 25 9.62 -2.07 3o /0.00 -2.00 .44 40 9.75 -1.86 50 8.98 -/.6/ L 60 765 -1.28 l

2.0 .40

O 70 6.39 - .95 l 462 - .64 U 90 254 - .38

k 95 1.38 - .25 /8 .36

O (-Q3) (' 13) C /.6 .32 L. E. Rod.: 1.58 m si oe orroe^s Ci through end of 26 ; p 1.4 .28 c chord: 4//S N O 24 2f /.2V .24 LID C.P.

w v q20p41 LOv.20^

u

/6' 61 t 0 0 o /2 ^ 81

60 /Ol .4^ .08

0 4m .2 .04 .o u a v 0 0 Airfoil: N.A.C.A. 4312 R.N.:3,/80.000 C Size: 5"x30" Ve/.(fl./sec.: 68.7

-4 R -.2

Pres.(s{i+d.otm):20.8 Dole:12-14-31 Where tested: L.MA.L. Test: V D.T. 731

-8 -.4

Corrected for tune/-wall effect -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.

22 REPORT NATIONAL ADVISORY COMMPPTEE FOR AERONAUTICS sto. UPS'. 4M, D t0 v /.25 3.32 - 60 0 u 2.5 4.47 -2.26

/I 44 v k - /0

5.0 6./3 -2.97 y o 7.5 737 -3.3/ t0 /0 0.36 -3.50 0 20 40 60 BO l5 9.85 -3.62 Pe r Cent of,or, 20 /0.60 -a60 25 11,32-3.55 36 301/. -3.50 .44 .09 4011.18 -3.33 l S 50/0. / -2.93

L

60 9.01 -2.42

2.0 .40 G.08 32

70 73/ -L66 529 -/.30

,v BO -c°

0 9 292 - .74

1.8 .36 x.07 28 i?

95 1.5B - .44 k

.d 100 (.l6) (-.16) 100 0 D C 1.6 .32 0.06 24, L. E. Rad.: 2.48 l W V u Slope afradius through end of 44 .28`c 00 .05 20 28 0 chord:4 /5 va

1.2 i.24X j.04

c. p. 24 0 20 C m o v L/D /. 0,E .20 4.03

420040

u

N ^ `c

.8 v./600 .02 e /6 60

o 'o

s °./2 .0/ 4^ 0

/2^8

.4".08 0 B o /0 0

e

-4o

.2 .04 G-./ 4

v 0 k m -8

0 0 0 -.2

z° 0 4

Airfoil., N.A.CA. 4315 R.1i:3,120,000

C

Size: 5"x30" Vel.(ft./sec.): 69.3 -4 R

-.2 w-.3

Airfoil: N.A.CA. 4315 Pres.(stnd. aim): 20.8 Dole:4-14-31

v

V Where Msled.•) A.L. Test.-VDT 565

Z-4

-.4

-8 Corrected for funnel-wall effect -6 -4 0 4 8 12 16 20 24 28 32 Lift coefficient, Q Angle of otfock, a (degrees) FIOREE 39. - N.A.C.A. 9315 Elrtoll.

Sto. Up 'r. L0, 0 W b0 0 - t 0 /.25 4.07 -/.90 2.5 S.36-2.74 y -/0 S. 7.20 -3.68 Z58.56-4.25 0 20 40 60 W /0 ) /0 9.64 -4.56 /5 //.22 -4.92 Per centofchord 20 /225 2512.

80 30 13.00 -500 .44 40 /264 -477 50 //.65 -4.24 D V$ 60 10.15 -3.55 2.0 .40 70 6.24 -2.76 60 5.95 -/.94 v 90 329 -/.OB 48 36

95 1.79 - .64 k

O 100 (./9) (-./9) too - o

a

/.6 .32 0 L. C. Nod.: 356 v Slope ofrod/us L;, Ci u Through end of 28 p 0 1.4 .28 o chord: 4 /5 va 24 20 1.2V.24w m c.p.

q 200040 40.^.20v a 0

LD N ^ u a4

o /6' 60

0 0

/2 o C 80 .6' .12

u

8 0100 .4'.08.

0 4v .2 .04 V`

C

.o u 0 0 y

4 04

Airfoil: N.A. C.A. 4318 R.N:3,09Q000

C U

Si- 5'k30" VeZ(ft/sec.): 69.5 -.2 Pres.(st^7datm.):20.8 Dofe:4-14-31 4 0.

v Where tested.-L.MA.L. 7-est.-VDT 566

-B u

-4 0 Corrected for tunnel-wall effect -8 -4 0 4 8 12 16 20 24 28 32 Angle of attack, a (degrees) Lift coefficient, C FtomE 35 -N.A.C.A. 4318 MO.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENBFPY WIND TUNNEL 23 Sto: ./2 0 o ut / 0 1.25 4.84 -2./9 129 l 25 -3./8 o -/0 44 50 6.26 - 4.38 R ,0/0. - 0 20 40 60 7 564 80 /C 15 1262 -6.19 Per cen/ofchord 2013.72 S. -42 /4.29 -6.50 25 h 30 1450 -6.50 .09 36 m .44 40 /4.09 -6.23 a 50 12.96 -5.56 D 60//. / -4.66 2.0 .40 V .08 32 O 70 9./8 -3.66 t 80 6.64 -258 N 90 3.67 -1.43 k 28 b 95 2.00 - .84 18 .36 O /00 (.221 (-.22) k D /00 - o 1.6 .32 0.06 4. L.E. Rod. 65 v s/ape ' C° 1 = , ihrou9 28 0 0 1.4 .28-, 0.05 20 w chord: 4115 C W ^ 24020 1.2V .24 x.04 /6 0 c.p.

a a '^ 20 40 /.0- .20 u x.03 LID N u a /6O 60 .8 v '16° .02 o O /2 °C80 .6X./2 .0/ 4y o u k vi 0 M 8'.100 .4".08 0 O'er C Y 4w -4 0 2 .04 G f u C ° 0 0 m-.2 - 81 ^° O 4 Airfoik NA.CA. 4321 R.N:3/20,000 C Size. 5"x30" Ve%(f/./sec.): 694 -4 0. -.2 -.3 Pres.(sfnd.ofm.):20.7 Dote:4-15-W c A/rfo/1.NA.CA 4321 R.N. 3,120, 0001 Where lested.-L.MA.L. Test: V.OT 567 Dole:4-15-3/ Test, VD.7.5671 8 u 4 1-4 Corrected for funnel-wo/1 effect Corrected to infinite ospectrobo ^ 0 .2 4 .6 8 /.0 /2 1.4 16 LB -8 -4 0 4 8 12 16 20 24 28 32 -4 Angle of oltock, a (degrees) L/ft coeff/'c/enl, C Ftovaa 38.-N.A . C.A. 9321 SIM% Up', L''wr .

sf. D /0 JS 1.25 -.64 u u 0 2.5 1.68 - 79 u,A 5.0 2.79 - .82 ass - O 40 60 80 /G i0 .60 20 /5 5.15 -.25 40 Per cent ofchord 20 5.90 .12 25 6.42 .46 30 6.76 .74 .44 .09 36 w 40 6.90 1./0 a so 6.55 1.24 D 60 5.95 1.27 2.0 .40 U.08 32 1.16 70 485 80 3.56 u N 90 1.96 .49 k 1.8 .36 ^.07 28 ti 95 1.05 24 /00 (.06) (- 06) .d 00 - o O 1.6 ,32 0.06 246 L. E. od.: 0.40 w l 51ape ofrodi 5 V through end of 20'.

2B o ( 1.4 .28- 0.05 chord: 4120 W C 24a 2( /.2V.24 v.04 L/D 168- c.p.

m a O 2004( lo-.20u x.03 12'^ u /6 6( .8 ^./6 0 .02 ^ t .6 0 ./2 .0/ 4s u k u 80/a .4".08 0 0Z O `^ C I .2 .04 0-./ -4o ° 4 u in C, ,o v a LIZ O O -.2 ^ 0 Q Airfoil.-N.A.CA.4406 R.N:3,10Q000 e c U 54, 5"x30" Vef.(ft/sec.): 69.4 -4 4 Pres.(sfd. nofm.):20.9 Dote:8-21-31 N.A.CA. 4406 R.N..:3,/00,000 U Where tested: L.MA.L. Test: VOT. 651 -21-31 Test: VO.T. 651 -B -.4 0-.4 Corrected for tunnel-wa//effect 8 -4 0 4 B 12 16 20 24 28 32 -4 Angle of oltock, a (degrees) L ift coeff/'c%ni. 0, F[0vae 37. - N.A.C.A. 4908 airfo.

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Sto. up-,. LWr.

a cu 2 48 u o 0 - O /.z5 /.e/ -LOS 0 2.5 2.61 -1.37 R o-/0 S. 0 3.74 - /.65 75 4.64 -1.74 0 0 20 /0 5.37 -/.73 40 60 BO / 0 40 /5 6.52 -155 Per centofchord 20 7.33 - /.30 25 7190 -L02 44 30 8.25 - .76 09 36 40.8.35- .35 50 767 - .07 D .08--e 60 7.00 .14 2.0 .40 at 32.

70 5.76 .26 60 4.21 .26 90 2.33 .14 1.8 .36 .^ .07 28 95 1.26 03 100 (.03) 1-.091 /00 0 a 1.6 .32 0 .06 L. E. Rod.: 0.69 v V u Slope o(rodius 'b through end of 20 1.4 .28 o 28 0 .05 c chord: 4 20 W a 24 a 20 .6 1 4 V ^ c.p.

Lm LO,v.20u 020040 .03 U y ^ 8.

r°

.$ v./s o p 16 60 .02

0 0 .6 U.12 4' 0 o /2.g 8 0/ w .4'.08 8 0/00 0 c .2 .04 t -4 ° 4 u

C

0 c

a -8 0 0 m 0 4 -.2

0 i

AirfOlL'NA.CA.4403 R.N:3,170,000 c Size: 5"x30" Ve1.(ft./sec.): 68.6 2 -4 0.

c Pros.(stndatirz): 2/.0 Dote:8-24-31 'WA .3 Airfoil. NA.C.A. 4409 R.N.:3,170,000 12 V Where tesfed L. .L. Test: V DT 652 _ 4 Dote: 8-24-31 Test: V. D. T. 652 -/6 -.4 -6 Corrected for tunnel-watt effect Corrected to inf'nite aspect ratio -4 -2 0 .2 .4 .6 .8 /.0 /.2 1.4 16 /.B 12 16 20 24 28 32 -8 -4 0 4 8 Angle of attack, a (degrees) Lift coeff/cteni,C Fmm. 38. - N.A.C.A. 4408 I& M.

S1o. Opi-. Lm, a c ^r / 0 1.25 2.44 -/.43 o - 0 2.5 3.39 -1.95 y 0 -/0 5.0 4.73 -2.49 h 7.5 5.76 -2.74 0 0 20 40 60 60 /0 /0 6.59 -2.66 /5 7.69 -2.66 Per centofchord 20 6.60 -2.74 25 9.41 -250 30 9.76 -226 .44 40 .460 -/.80 50 919 -1.40 60 8.14 -L00 2.0 .40 70 669 -.65 80 4.89 - .39 so 271 - .zz u /.8 .36 k 95 /.47 - ./6 C l00 1.131 (-./31 100 - O a 1.6 .32 c L.E. Rod.: 1.56 m V 5/ape of-d,11, 0 lhrough endpf 1.4 .28 chord: 4 20 W A2V.24 24 1 - 20 c.p.

LI D w v LO y.20° Q20 a4 0 U u l $ .!6 ri 16 6 0 0 0 .6'./2 0 12 8 .4'.08 80/0 0 `^ C .2 .04 ° 4 u C .0 0 0 R° 0 4 Airfoil, N.A.CA. 4412 R.N:3,200,000 C Size: 5"x30" Vel.(ft./sec): 66.6 _4 0.

-.2 Pres.(sfnd.oim.):20.8 Doie:12-15-31 Where tested: L.h1A.L. Test: V D.T. 732 _g v -.4 Corrected for tunnel-wail effect - -8 -4 0 4 8 12 16 20 24 28 32 Lift coefficient. 0.

Angle of attack, or (degrees) .-N.A.C.A. 4412 91doil.

1NOME 30 CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 25 S0. UP r. L'0r.

^D l0 U /.ZS 3.07 -1.79 0 2.5 4./7 -2.48 l U 44 -l0 5.0 5.74 -3.27 4 p 75 6.91 -3.7/ /0 7.84 -3.98 0 20 40 60 BO /0 ./0 15 9.27 -4./8 Per cent ofchard 20 /0.25 -415 - 25 /0.92 -3.9B .09 36 30 //.25 -3.75 .44 //.25 -3.25 5 /0.53 -2.72 D 60 9.30 -214 °° 32.

.08 2.0 .40 U 70 7.63 -/.55 O 80 5.55 -/.03 _6 90 308 - .57 U .07 LB .36 ._6 95 1.67 - .36 k O 00 (.16) (-.16) Lc .06 24'

/.6 .32 0

C L. C. Rad.: 2.48 W slope o7u9 V u

i 0

through end o7

^.0$ 1.4 .28 o 28

Chord: 4120 W ^ 3 .

04 /.2 V . 24k m 24 '

e.p. a 20

.03 2 /.0.200 q20040

L/D a 8'

.02 A./6^

0 16--60

0 0 oe .0/ 4' .6u .l2

0 0 .4'.08

x 80/0 0 `• C 0 4 v .2 .04 c

.o u

-.2 -8 0 0 m

a 0 0 4

Airfoil: NA.C.A.44/5 R.N:3 /10.000 .c Si- S"x30" Vel.(fl/sec.): 69.5 Airfoil. NA.CA . 4415 R.M.: 3,110.000 /2 P(sti7d.otm):20.7 Dote:8-26-31 c

v

Zr.

_ Dote:8-26-31 Test: l!D.T. 654 Whe tested: L.MA.L. Test: VOT. 654 _4 -16

0 -8 u

' 4 Corrected to in fihite aspect rofo Corrected for tunnel-wall effect -4 720 .2 .4 .6 .B LO L2 /.4 L6 /.8 -8 -4 0 4 8 • 12 16 20 24 28 32 Lift coeff can't C Angle of attack, a (degrees) FIGURE 40. - N. A.C.A. 4415 airfoil.

r. Lm, 5/a. Up • o l0 us ° - ° 0 /.25 376 -2// 2.5 Soo -299 W k _/0 5.0 6.75 -4.06 C o 7.5 6.06 -4.67 0 20 40 60 80 /0 /0 9.// -506 0 /0 15 10.66 -5.49 Per centofchord 20 //.72 -556 h 25 12.40 -5.49 44 .09 30 12.76 -5.26 36 w 40 12.70 -4.70 a s0 /1.65 -402

a

3. 60 10.44 -24 32 v 2.0 .40 ^ 08 70 6.55 -2,45 O 80 6.22 -/.67 N 90 346 - .93 AS .36 ^.07 28 Z{ k 95 /.89 - .55

g

24 0 1.6 .32 0.06 C 4.E.Rad.:3.56 If l W G u Slope ofrod)u5 Z' D through endof 20 m /.4 .08- 0.05 28 0 chord 1.2V .24 •F .04 2 v 24 0 c. .

a v v `^ 2.'c 40.y .200 ^.03 q 2004 0 k .0 N ^ 0 L/D .8v.16^ 02- /6^ 6 0 0 4^ 6 u .12 .0/ 0 o /2r8 k u 0 Oa .4 08 8 0/0 0 3 `• C C ` 0 4 v -4 o . 2 .04 u n - B 0 0 m -.2

0 4

Airfoil. N.A.CA. 4418 R.N.:3,100.000 c Size: 5"x30" ve/(f//sec.): 69.5 -/2 -.2 u-.3 Dcle:8-27-3/ I: NA.CA. 4418 R.N.: 3,/00,000 I-rl' v Where tested: L.MA.L. Test: l!0.T 655 U 9-27-31 Test: VD. T. 655 -16 0 -.4 -.4 Corrected for tunnel-wall effect -8 cted to infinite aspect ratio .4 .6 .8 LO L2 1.4 /.6 /.8 -8 -4 0 4 8 12 16 20 24 28 32 .4 Angle of attack, a (degrees) Lift coefficient, C FIGURE 41.-N.A.C.A. 4418 airfoil.

270770-35-4 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS - zo L^ r D ./2 48 4.45 -2.42 g 0 2.5 5.B4 -3.46 w _/0 W 44 50 7.82 -4.78 p h ]5 924 -5.62 0 20 60 l /B %35 -6.15 40 80 00 ./0 15 12.04 -6.75 Per cent ofchord r 2013.17 -6.9B.

n y 25 13 B6 -6.92 30 14.27 -s.76 .44 .09-- -36 v 4014.16 -6./6 8 ^ 50 13. lB -5.34 D BO //.60 -4.40 2.0 .40V .08 a; 32 U 9.50 -•235

t 60 6.9/ -23/

N 90 3B5 -/.27 1.8 .36 u.07 26 e `o 100 /.- F0 O l0 0 0 1.6 .32 0.06 24 0 LG Bod.: 4.65 v ^ V U Slope ofradius through end of 28 0 C /4 .28'^ 005 20 u chord: 4/20 v ^ Q 241,20 /.2V.24 y v.04 168 .

c. p r-

m 12 Q 20x40 /.'O w .20 u °.03 R y ^ U ^ o /6^ 60 B v . /s 0 .02 L/D o b ° 0 u 0/280 .6 ./2 .0/ S uk 8 0/00 .4".08 0 0 z ° 4 u C 0 Q 0 0 -.2 -8 u' Airfoil NA.CA. 442/ R.11:3.110,000 e C u Size: 5"x30" Ve/(H/sec.): 69.4 -4 0. .2 /2 Pres.(stird ofm): 20B Dote:8-28-31 3 Air foil , N. A. A. 4421 R.N:3,110,000 w ' Where tested: L.M. A.L. Test: VD.T 656 Dote. B-28-31 Test: VD.T. 656 -16 -B -.4•4 Corrected for tunnel-wall effect Corrected to infinite aspectrotio R- 20 24 28 32 -4 72 0 .2 .4 .6 .8 LO /.2 1.4 -8 -4 0 4 8 12 16 Ang le of attack, a (degrees) Lift coefficlent,C FIGURE 42.-N.A.C.A. 4421 airfoil.

5f. UPr.G' r.

D 25 /. -. UC 7/ 0 ' .66 U _/0 e 5.0 260 -/.00 p h 7.5 3.25 - .97 2.5 /0 362 -.89 20 40 60 80 00 /5 4.74 -.64 Per centofchord 40 20545 -:32 25 59B .02 30 6.36 .34 36-0 .44 .09 40 6.74 ,93 50 6.65 /.35 0 0^ 60 613 /.56 .40 ro 2.0 -08 32 O 0 70 5.21 /.53 60 390 /.25 U 2 90 /8 .72 1.8 .36 .07 95 1.17 28 /-0sl /- i°o°o /.6 .32^ x.06 24:9 L.E. Rod.: 0.40 W 51ope fro od us through entl of 28 p 0 44 .28 , 0.05 '01 chord: 4125 y N 24 0 20 /.2V.24^ y.04 /60 L D C k k c. p.

Q 20F40 /.0.1.200 0.03 12' u uo V ^ u g` /6' t 60 .81./60 .02 ,o 12- BO .6 u .12^ .0/ a w 4u 8,0/00 .4 .08 0 0 k c r ° 20 4 u .o 0 v 0 Q O O y -.2 -8 ^` Airfoil: NA.C.A. 4506 R.N,:$05Q000 e c °u Size: S"x30" Ve%(H./sec.) 70.0 -4 0.

Pres.(st'n: d o tm.): 20.6 Dafe: 4-/5-31 c •3 •2 Airfo#-N..A.C.A. 4506 R.N:3.050,000 -l2 _8 d M. Where fesfed• L.A.L. Test: V.O.T.568 Oote: 4-/5-3/ Test: V AT 568 _l6 -.4 - 4 Corrected for tunne/-/l effect 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 27 up', V0, 51.

D /O U ^ /.25 13S -1.12 0 2.5 2.47 -1.50 L, -/0 -1.84 5.0 3.54 4.36 -1.99 7.5 0 20 40 60 80 /0 /0 SOS -2.05 Per cent ofcho^d /5 6.12 -/.96 -1.75 20 6.89 25 7.46 -/.47 30 7.85 -1.16 .44 - .52 ,40 6./9 50 7.97 .03

a

60 7.27 .42 2.0 .40 70 6./2 .62 80 4.56 .60 90 256 .35 U /.e .36 k 95 /. /5 100 (.09) f-.091 /00 - 0 /.6 .32 C L. E. Rod.: 0.69 v afrad'us V Slope through end of /.4 .28c 28O chord:4 25 ,W 24 2i /.2V .24 LID 0 w v N 20041 LOv.20^

p /6 61 8x./60

o O D o i °12^Bi .6h ./2 .4'.08 8 0 /Oi c C .2 .04 ° Su o v O 0 ^ 0 Q Airfoil, N.A.C.A. 4509 R.N..:3,120.000 .0 Size: 5"x30" V.I.(fG/sec.): 69./ -.2 -4 0.

P es.(s fnd. almf:20.9 Dote: 4-15-31 Where fested: L.MA.L. Test: V D.T. 569 G -.4 Corrected for tunnel-walleffect -B -8 -4 0 4 8 12 16 20 24 28 32 Lift coefffcient,Q Angle of attack, a (degrees) Proves 44.-N.A.C.A. 4608 airfoil.

L'-r.

5o.

Up', vD /0 125 2.33 O 25 3.22 -2.07 O SO 4.50 -2.67 7.5 5.46 -2.99 /0 6.25 -3.18 0 20 40 60 60 15 7.46 -3.28 Per cont ofchord 20 a34 -3./7 25 8.95 -2.95 9.37 -266 .44 40 9.64 -/ -/.97 50 9.29 .29

U

606.43 - .70 l 2.0 .40 70 7.06 -.2,9 ° -:04 t 805.23 U 90 2.93 .00 48 .36 /.58 - .04 /00 (./2) (-. 12) C

a /00 - 0

1.6 .32 C LX. Rod: 1.56 v Slope ofrodus Ci through a 2Z 28 0 /.4 .281 chord: 4/25 W 24 L. 2 2,j. 24 i?

C.P.

LID ^o X20041 /.0 r .20 0 N ^ u o /6^ 61 .6 .160 r o /2 C e .6°./2 u k B o /Oi 4".08 0 v 4 .2 .04 u o v ^ O4 0 0 AlrfOlL'NA.CA.45/2 R.N:320Q000.

C Size: 5'k30" Ve1.(fL/sec): 68.2 -.2 4 0. Pres.(sfnd ofm):2/./ Dafe:/2-/6-3/ v Where tesfed'LAIA.L. Test:V.DT -8 -.4 Corrected for tunnel-wall effect - 0 - 4 O 4 8 12 /6 20 24 28 32 Angle of attack a (degrees) Lift coefficient,C F'lovac 46 .-N.A.C.A. 4512 airfoil.

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS St.. Up". L'w^ - .

/0 .12

48 w o

0 - o

u ^

o /.25 2.94 -/.88B

2. 5 00 -2464 44 4 0 4 49 5. -3 8 0. o 7.5 6.60 -398 7 to -4.30 0 20 40 BO 10 ].4s 60 IS 8.85 -4.59 40 Per cen/ofchord 20 9.6/ -450 25 /0.47 N 3010.80 -4.16 36 w .44 .09 4011.019 -3.44 501062 -262

D ti

a 60 9.56 -1.e5 32 O 2.0 .40 •x.08 70 299 -1.20 BO SB9 s° - .69

v

90 332 - .35

28 ZS 48 .36 x.07

O Ci .d 100 - O L6 .32 x.06

24 p c

L. E. Rod.: 246 N

Slope o/ radius v

I= 9,5 end of 2B 0 /.4 .28- 0.05 20 u chord: 4125 o va 24 20 1.2V .24^ v.04 /6g C.P.

ti v v y 20040 /.O.w.20' x.03 12 1 LID J U l 1 h /6^ 60 6 ^^ .B m'/6^ .02 o '0 k o /2c80 .6 u .12 .0/ 4y w 8,0100 .4J .08 O

4 0

C

4 0 v

-4 0 .2 .04 V-./

C .o u v O 0 w -.2 -B 04 Airfoi/-NA.CA.45/5 R.N.:3,080,000 C U Size: 5"x30" Vel tft./sec.): 69.6 -4 R P es.(sfndofm.):20.7 Dote:4-/6-3/ v Where tested'L.MA.L. Test: VD.T. 571

-8 0-.4

-.4

Corrected for funnel-wall effect -8 -4 0 4 8 12 /6 20 24 28 32 -4 -20 .2 .4 .6 .8 /.0 Ang/e of Mock, a (degrees) Lift coefficient,C Fca y sE 46.-N.A.C.A. 4518 airfoil.

'r. Slo. UP LWr a cu /0 y 0 - 0 a /.25 3.56 -2.25 l 0 25 4.79 -316 0 0 6.49 -4427 .// 80 IC /0 6.74 -5.4/ 0 20 40 60 I /5/0.25 -569 Per centafchord 40 20 //.27 -6.01 2511.96 30 12.37 -5. ' 9 ' -5.66 44 .09 - - 36 w 40 /253 -4.69 l 50 1/.94 -3.94 60 /0.72 -296 ..._.......

2.0 .40 .x.08--_.._ 32 a

70 6.92 311

s 60 657 -1.34

v 90 3 69 - .7/ Cno

k 95 2.01 - .44 /B .36 y.07 28 ti

/00 (.19) (-./9) k

a /00 - 0

1.6 .32 0.06

L. E. R. d.: 3.56 d.

v Slope of-d,,,

C

26 0 1.4

0 .05

c hord 4125of .28`c W a

24 , b 20

L2^.24X v.04 168

c.p.

O 20 0 40

1.0 .20 4.03 .a

y ^ k

U IL /6^ 60

.8v./6o 02 6i

o /280 .6 .12 0 .0/ a' 4y u 60/o0 .4 .08

4 0

`^ C 0 4v .2 .04 V`-./ o u C m 0 0 m -.2 Airfoil N.A.CA. 45/6 R.N:3130.000 0 e c • Size: 5"x30" 3 e).(ff./sec.): 66.9 -4 0. .2 .3 Pres. (sthd afm): 2L0 Dote:4 Airfoi/: NA.CA. 4518 R.N: 3 /30.000 Wherefested.-L.MA.L. Test.VOT 572 _ _ Dote: 4-16-31 Test, VD. T. 572 -8 u .4 .4 Corrected for tunnel-wc , H effect Corrected fo inflhVie ospectrofio -8 -4 0 4 8 12 16 20 24 28 32 -4 -2 0 .2 .4 .6 .8 /.0 L2 14 /.6 L Angle of otfoch, a (degrees) Lift coefficient, 0 Frooxx 47 .-N.A.C.A. 4518 airfoil.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE -DENSITY WIND TUNNEL sl. . L'v, r.

UP' r U /O /2 u 1.25 4.23 -2.56 i 0 5.60 2S -3.B6 N k -/O SO 750 -505 p .//

94Mm

7.5 0.90 -5.90 /0.00 0 20 40 60 7 /0 -6.50 80 /C E V 1511.63 -7.17 Per cent .{chord co ./0 40 20 1273 -7.42 P 2513.46 - 7.38 30'3.06 - Z/6 .44 .09 311 40 1397 -635 60 13.27

50 -5.27 l

11.86 -4. /2 2.0 .40 h .08 32 D 70 9.65 -3.00 o, 80 7.26 -1.97 S W 90 406 - /.05 L8 .36 o .07 95 2.24 - .63 28 tt /00 (22) (-22) 100 - 0 1.6 .32 0 06 24 p LE.Rod.: 4.85 m Ci 5/opeofrodius l Ci u through end of 26 O 1.4 .2Bc 0 .05- zo chord: 4/25 0 v v $ 24 t. 20 L2 V .24 v .04 p 16 c. p.

k v 0 Q20E 40 LO .03 o y u 16 -60 ev.160 02 ° B L/D L 0 o O 0 0/2280 .6^./2 0/ qt u 60100 .4 .08 0 c 2 .04 -40 0 4u C p Co v _2 0 Q 0 0 y -80 Airfoi/: NA.C.A. 452/ R.N.:3,150,000 e c ou 6/ie: S°x30" Ve/.(ft./sec.J: 69.0 -4 .2 .3 Pres.(sfnd.otm).•20.6 Dote: 4-/7-3/ c Airfoi..• N.A.CA. 4521 R.N3 : /50,000 /2 Where tested:L.MA.L. Test: V.D.T.573 Dote: 4-17-31 _4 Test: VD.T 573 _16 -6 -.4 Corrected for tunnel-wall effect 0 Corrected to infinite ospect rotib -6 -4 0 4 8 /2 16 20 24 28 32 .4 .2 0 .2 .4 .6 .8 LO 1.2 /.4 t.6 L8 Angle of otlock, a (degrees) Lift coeff/cient,0 FIGURE 48.-N.A.C.A. 45 toll.

S^. U p ^. Gwr.

ITT QD /0 /.25 -1.57 2.26 i 0 2.5 3./3 -2./6 //j - 4 -/O SO 4.36 -2.81 7,5 S. -3.17 O 40 60 /C /0 6.02 -3.40 20 60 15 7.17 -3.56 Per cent ofchord X140 20 -3.50 6.0/ 25 8.60 -331 ,44 30 90l -3.00 40 936 -2.27 50 9.16 -/.4/ O 60 6.56 - .56 2.0 .40 70 7.44 ./0 C L 60 5.66 .40 N 90 323 .3/ /.8 .36 .0 0 k !./2) 100 (-.%21 /o0 - 0 a /.6 .32 0 +-H 24 ROd :/7 _ 6 T_ 1-T v V u 5/ape cf "of through end of 1.4 .28 C 0 28 0 chord: 4/30 N L2V .24w v 24 0 20 16 c.p.

v '^ L/D /.0_v .20 u 20 a 40

u p, 4

.Bo .lso g16D60 .6' ./2 o /280 w .4'.08 0 8 0 /00 .2 .04 VE 4 u J o u a 0 0 y 0 Q 0 Airfoi/.• N.A.C.A.46/2 R.N.:3,2/0,000 c Si 5"x30" Vel.(fl./sec):68.3 -.2 -4 0.

P es.(slnd.olro):21,0 Dole: /2-/7-3/ Airfoi/:NA.C.A. 4612 R:N:3,210,000 Where tested.-LM A.L. Test: V.D.T. 734 Dote: 12-17-31 Test: V D. T. 734 -l( _4 0

- 8 u

Corrected for tunnel-wall effect Corrected to 1/7fhffe aspect rotio -8 -4 0 4 B l2 /6 20 24 2B 32 -4 Angle of ottock, a (degrees) Lift coeff/cie,L Cc F-it. 49.-N . A.C.A 4612 airfoil.

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS St . up L-- No /0 /.25 221 -/.6/ U t 0 253.04 -222 k -/0 50 4.24 -292 y ° 44 10 5.13 -3.32 O 20 40 60 80 /L l0 5.64 -3.56 /56.95-3. 79 Per ceniofchord 0 20774 -377 256.3/ -3.60 30 9.71 -156 09 36 40 9.06 -2.56 50 6. - /.64 60 6.47 47 - .65 2.0 .40 U o .08 70].6] .34 0 BO 6.21 .95 N 90 3.73 .78 U 18 .36 .0 .07 28 95 .42 02) Ci 100 L/2) (- 0 /00 0 16 .32 0 .06 24 L. E. Rod.: 1.56 v V U Slope ofrodLs ihraugh4/3s endof /4 .28c 28 D 0 °.OS 20 chord.

v ^ A2,j. 24 b 24 0 20 .04-/6 v '^ /.Oy.20° 20 40 .03 /2 0 ao u ^4 l .02 p 16 ^ 60 8 0 L 0 .12 .0/ 4 .6 o / t 80 k .4 .08 6 °loo 0 0 40 .2 .04 y` -4 U n .0 v 0 0 y -8 Airfoil: NA.C.A. 4712 R.N:3,160.000 u C Si-S"x30" 1/e/(tf./sec): 68.7 2 .3 Pres. (sind.oim):210 Dote:12-18-31 c 0.

Z- toil. IV. A.0.A. 4712 R.N.: 3, 160, 000 /2 Where fested'L.MA.L. Test: V.D.T.735 _ Dote: 12-18-31 rest: VD. T. 735 E _.4 -8 lB Correctedfor iunnel-wo//effect ' 4 Corrected to infhfte aspect ratio -8 -4 0 4 B /2 /6 20 24 20 32 £ -2 0 .2 .4 .R A CO 12 14 1R 4R Angle of ottoc/, or (degrees) Lift coe fficienf, C FIGURE 50.-N.A.C.A. 4712 airfoil.

Sfa. up'r. L'0, O cu /O y ./2 48 0 - 0 U l 0 /.2S 370- .79 2.5 4.99 - .94 y M. 10 44 50 6.92 - ,79 y p 7.5 6.42 - .46 /0 9.56- ,15 0 20 40 60 80 10 7 /s //.09 Per centofchoro 0 20 //.74 .26 25 //.92 .03 3011 Co, ; - . /O .44 .09--36 4011. -./7 50 /0.49 :.12 60 9.// .05 32, 2.0 .40 G 0 .08 70 7.37 .0/ C 80 5.30 .02 °C 90 2.99 - ,03 95 1.55 -: (L /.8 .36 .07 O loo (,/2) (-.", /00 - O a 1.6 .32 .06 24' L E. Rad.: 15B v 5lope 3 ofrodi s V ihro 3 9h end of 28 p 0 1.4 .28 0 °.O5 20 chard: 6110 N ^ 24 kk 20 L2V.24- v 04 /6 c.

^o v LID.

20 40 40i .20 o .03 /2 h V u a o /6. 60 .8 tE ./6 .02 8 o 12,80 .6u .0/ 4 8 0 /00 .4". 06 0 0 n .2 .04 Ci -41 ° 4u _.2 _6 0 0 y 0 4 Airfoil: N.A.C.A. 6212 R.N.:,^240,000 c u Size: S"x30" Vel.(ft./sec.): 66. l -4 0. .3 Pres.(st'nd. ofm):20.9 Dote:12-21-31 Airfoil: N.A.CA. 6212 R.N. 3,240.000 -/2 Where tested: L.M.A. L. Test: V.O.T. 737 _4 Dote: /2-21-3/Test: V D.T 737 _16 -8 -4- Corrected for funnel-wall effect 0 Corrected to infinite ospecf ratio -8 -4 0 4 8 12 /6 20 24 28 32 -4 -2 0 .2 .4 .6 .8 LO l.2 /.4 1.6 c` /.8 Angle of offock, a (degrees) Lift coefficient, C FIGURE 51 .-N.A.C.A. 6212 airfoil.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 31 sf.. L r.

UP r v ^ /O 125 1.63 - .40 u 2.5 246 -.34 l -/O 50 3.79 .OS y 0 75 4.67 .50 to SB0 .96 0 20 40 60 80 1 00 /5 7.26 /.79 Per centofchoro ./O 8.23 2.45 25 8.80 2.64 y 30 9.00 3.00 .44 .09 36 40 6.76 2.98 50 6.17 2.86 60 720 2.62 2.0 .40 x'-.08 32 70 5.90 2.21 80 4.27 1,63 u 90 2.34 .86 95 1.24 48 .36 ^.07 28 tt .43 k 100 (06) (-.06) l00 - O 0 1.6 .32 0'.06-- L.E.Rod.: 0.40 v 5/ape of'oo- G U lhra 3 qh end of 28 0 0 1.4 .2B 0.05 20 w chord: 6//5 c .v 24 ^ k 20 /2V.24-Z^ u v.04 /6 0 C.P.

LID v `^ q 20040 40y.20, k'.03 12' k u O, 0 /6 60 .8v.16o .02 o a k /2U 80 .6.12 .0/ 4^ w B 0/00 .4 J .08 0 0p g C 0 4 W .2 .04 -4 0 .o u ro v 0 O -.2 -8 Airfoil: N.A.C.A. 6306 R.N:,$080,000 C Size: S"x30" Ile%(fr/secf:69.8 -4 0.

-.2 -.3 P es. (st'nd. otm.):20.6 Dofe: 4-/7-3/ c AirfOi/:NA.CA. 6306 R.N.:3,080,000 L.M.A.L.

Where tested. • Tesf: VD.T. 575 Dote:4-17-31 Test: V D.T. 575 -8 -.4-.4 Corrected for tunne%wol/effect Corrected to infinite aspect ratio 4 -, -8 -4 0 8 /2 16 20 24 28 32 -4 O .2 .4 .6 .8 /.0 /,2 44 /.6 /.

Ang/e of aHo'k. a (degrees) Lift coefficient,C FIGURE 62,-N.A.C.A. 6306 airfoil.

Slo.

/O O U /.25 -.75 5.0 r 0S 76 7.5 50 0204060BO/O /0 ./B /5 .46 Per Gent ofchord 0 ./0 40 9.70 1.02 25 10.29 1.35 h 30 10.50 1.50 .44 .09 360 40 10.24 1.53 8 l 50 9.50 1.55 60 6 .35 /.4B 2.0 .40f 70 633 /29 S 60 90 4.94 . 50 W 90 2.71 .50 L8 .36 •u.07 U 95 28 tl /.45 23 /00 cos) !-0 C -

a loo o

/.6 .32 '0.06 L.E. Rod.: 0.89 0 u N Shope ofrad'us /hrough end of 1.4 .28-, p.05 28 0 chord.' 15

v$ 3

/. 2V.24 ,4y v.04 24 0 20 160 C.P.

L/D v '^ v /Ov.20a x.03 q20 40 c U ^^ .02 p 16 D 60 6C l l O ^ O .6i .12 .0/ o /2 ^ 80 4^ o u .4^ .08 0 80)00 0:,L, .2 .04 V -./ 0 4 u -4 0 C v v 0 0 y -.2 ^ OQ Airfoil: NA.CA.6309 R.h!:3.110,000 O C V Size: 5"x30" Vel.(R./sec.): 69.4 -.2 ^-.3 40. Pres.(sihd.olmJ:20.B Dole:4-18-31 U Where fested: L.MA.L. Test: V.D.T.576 4-18-31 Test: V. T. 576 -.4 0 -.4 -8 Correc led for funnel-wo//effect Gted to fnfnile aspect ratio -8 -4 0 4 B 12 16 20 24 28 32 `C -.4 .4 ".6 .8 /.0 1.2 /.4 16 1 Angle of otlock a (degrees) Lift coeffi'cient,C FIUVSE 63.-N.A.C.A. 6308 airfoil.

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS St... Up. L'w'r. 'r 1b /0

./2 48

-6 --- /.25 3.05 0 U o x 25 4.20 -1.38 est No. 738 yk y0 .// 44 5.0 5.93 -1.56 oo q 4 0 Ll 7.26 - /.47 40 / /0 9.36 -/. 29 0 20 60 80 00 do 40 15 /0.03 - .83 Per cenl of chord 20 //. 14 -40 25 /1.79 - . 14 30 /2.00 .00 44 09 36 40 11.69 .08 50 10.84 .23 D G$ 60 9.50 .35 32, 2.0 .40 .08 70 7.76 .39 O 90 5.62 .34 W 90 3.08 ./S 1.8 .36 w .07 95 1.66 .03 /00 o0 l00 (./2) (-./?j C a 1.6 .32 0 06 24' G. C. Rd.: 158 U W S/ape ofrod'us yn endof 1.4 .2,9-, .05 20' 28 0 0 chord: 6/I5 /6 /.2V.24' w .04 24^ 20 C.p.

L/D v '^ 40 .208 o .03 /2: 200 40 q 8' e`* ./s o .02 o /sa 60 0 0 4' .6X ./2 .O/

o /2 8 0

u

w u

8'-/00 .4J.08 0

c 0 4v

Ile -4'

.2 .04 G U C O ll C.

v _2 -8 0 0 w Airfoil: N.A.C.A. 6312 R.N.:3.f80,000 i ou Ve1.(R../sec.): 68.5 size: 5"x30" -4 0.

.2 Pres.(sfnd.otm):2/.0 Dote:12-22-31 c R.N.:3,180,000 -/2 ' 3 Airfoil: N.A.CA. 6312 v -8 u

Where lested: L.MA.L. Test: V.D.T. 738 _ Dole: /2-22 3/ Test: V. 0. 738

-/B -.4

Corrected for tunnel-wa//effect 0 '4 Corrected to infinile ospect rolio

-4 -2 n .2 .4 .6 .,9 LO L2 14 L6 1.8 -6 -4 O 4 8 12 /6 20 24 28 32 F Angle of attack, o: (degrees) Lift coefficient.0 FIOVAB 64.-N.A.C.A. e: W.H.

Sic. up'r. LM, U /0 4 0 /25 399 /.84 25 5.15 - _/0 t0 SO ].OS -2.27 75 948 -2.39 "3 20 40 60 BO 10 /0 9.6] -236 0 i ,fcha d /5 //.45 -2.13 Per cen 20 /2.60 -1.63 25 /3.25 - 1.62 .44 30 /3.50 -/.50 40 13.15 -/.37 5012.16- U§ U .77 32 2.0 .40 60 r0.67 - o° 0 ]O 8.7/ -.50 c BO 6.30 - .30 90 a45 - .20 28

/.8 .36

'w /.BJ-./J /00 (./6J (- 0 /00 - 24

/.6 .32 0 C

L.E. Rod.: 248 V u v 5/ape ofrod'us 0 through endof 20

1.4 .28 o

28 0 chord.' 6//5 v /6 1.2x.24 24 0 2 0 c. p.

k k v 1.0 ,v .20

020;4 o

L/D U 6^./6 0

^ p /6 6 D

O O l l a ° 4 .6X./2 0 l2r 8 o

k u

.4 .08 8 0 /0

0 4 v

.2 .04 G C o u 0 0 m -8 Airfoil. N.A. C.A. 6315 R.N:3,100,000 U

C

Size: 5"x30" 3 el(0.1sec.): 69.7 /2 -4 4 c _ Pres. (surd almJ:20.5 Dote:4-20-31 Airfoil. NA.CA . 6315 R. N: 3100,000 Dole: 4-20-3/ Test: V D.T. 578 -16 Where tesled.- L.M.A.L. Test: V DT 578

F _4

_8 v _ c Corrected to inf nice ospect ratio Corrected for tunnel-wo/l effect 2 4 R R /0 /2 14 /R IR -8 -4 0 4 8 12 16 20 24 28 32 -4 Lift coefficient. C Angle of allocic, w (degrees) FIGURE 56. - N.A.C.A. 8316.11U1.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 33

w o /0 u ^ O o -/0 0 20 4060 60 /00 Per cent ofchord 1 ./0 h .44 .09 'z.e2 v O 1.40 U 2.0 .40 ..08 e 4 5 5 D v V LB .36 .36

a

-./9l 0

D 16 .32 0.06 .5a v U Y'us C 1.4 .28 f 28 0 0 ,05 my 24 a L2V .24'M x.04 160 c.R k v `^ v l.Ov,208 11.03-- q 20 0 l2c 4 k

U i

L/D .8^.16^ .02 0 16 l 0 0 6X .12 .0/ 4y 0 /2 k U .4' .08 0 03 90 Y n 0 4 N 2 .04 -4 0 .0 U v k _ W 0 O U 2 D C Airfoi/: NAC.A.6318 R.N:3080,000 C Size:5"z30" V.Z(ft/sec.): 69.3 _ ' 2 .3 _/2e P es. (sfnd.otm):2/.0 -40.

Dofe:4-20-3/ Where tested: LMA.L. Test: VD.7.579 Dole: 4-20-31 rest: V. D. T. 579 --4 .4 .4 Bu Corrected for funnel-//effect Corrected to k7finite aspect ratio l -.4 48 12 16 20 24 28 32 .2 .4 .6 .8 l.0 1.2 1.4 1.6 1.8 _?91' of attack. a (degrees' Lift coe ffcienf,C 56.-N.A.C.A. 6318 airfoll. Fcccaa Ito. Up'r. L' O wr.

1, o /0 /.25 5.70 -1.79 u S O 2.5 7.20 -2.65 l U 0 -/0 .// 5.0 9.36 -363 7.5 11.03 -4./5 0 20 40 60 80 /0 l0 12.35 -444 l5 14.26 -454 Per cent ofchord 20 15.54 -4.65 -4.58 25 /624 h 30 16.50 -450 36% .44 .05 40 16.06 -4.25 50 14.64 -3.7/ l

D G$

60 12.99 -3.02 2.0 .40 •.06 32 D 70 /060 -230 80 7.69 -1.60 v 90 4.22 - .90 1.8 .36 . 0.07 28 21 ti 95 2.31 -.55 ( 22' ( 02) d %00 a 1.6 .32 0.06 24 C L. E. Rod.: 4.85 's l 5/ape ofrad LS through end of C I 1.4 .28-, 0.05 20-.

28 0 chord: 6//5 v va 24 2 420.240 m.04 /6 p wa cp m `^ v 10y.20 q 20x4 .03 i 2c y ^ .0 .BV./6o .02 8c o 16D6 LID C 0 0 0 ^ .6,U ` .12 .0/ 4^ o /2 B 8010 .4' .08 0 o^ .2 .04 -4 0 : o u C w C v 0 Q 0 0 -.2 -8", Airfoil.•IV.A.CA.6321 R./V.:3/40.000 0 Q C u Si- 5'x30'Uel(ff./sec.): 69./ -4 0. -.2 ^ -.3 Pres. (stird.otm):20B Dof-4-21-31 R.N:3,140,000 AirfoiLNA.CA.6321 v Wherefesd:LMA.L. rest: V. D.r5B0 fe Dote: 4-21-31 7ez f: V. D.T 580 1_ -8 -.4 0-.4 Corrected for tunnel-wol/effect Corrected to infinite aspect ratio FTT - R 4 -2 0 .2 .4 .6 .8 LO 1.2 /.4 l.6 1.8 -8 -4 0 4 6 12 16 20 24 2B 32 -.

Lift coefficient C Angle of atfacl, a (degrees) FrooaE 57.-N.A.C.A. 6321 airfoil.

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS S e. u,. - ,. L'W^.

zD /0 .12 Z25 /.45 -.52 0 2.5 2/6 -.55 L,4 SO 3.32 -.36 y p-/0 7.5 4.24 -.06 /6 /0 506.2B O 20 40 60 80 /5 &J9.97 Per centofchord ./0 7.42 1.61 25 8.16 2./6 30 8.64 2.62 36 44 09 40 6.90 3.10 50 9.49 3.19 U$ 60 7.64 3.05 2.0 .40 .0B 32.

70 6.35 2.66

c

ao ass 2.oz sa 2.59 /. u .07 28 w /.8 .36 95 /.3B 55 /00 (.061 (-.061 o 100 - 1.6 .32 0 06 24' L.E. Rod.: 0.40 U W Slope If-d",C lhrough

28 p 0 /.4 .28- 0 .0$ 20

chord: 6/20 •v a 041, 20 1.2V .24 .04 /6

w

v '^ q 20 o40 LOV.20u^ .03 /2' 0 8

/6160 A ./6 0 .02

80 .6U./2 O/ 4' u 4^ 12 S 4^ .08 8 0 /00 0 0: kc .2 .04 C

° 4 u

C .o ° u

0 O o

0 4

Airfoil: N.A.C.A. 6406 R.N.:3,100,000 i c u Size: 5"x30•' Vel. (ft./sec.J: 69.2

-.2

Pres. (stud. ofm.): 2/.0 'Dote: 8-31-31 -4 0. Airfoi/: N.A.CA. 6406 R.N: $ /00,000 -/2

v

Where tested.• L.hLA.L. Test V.O.T. 65B _ 4 Dote: 8-31-31 Test: U. D. T. 658 _16

-.4

Corrected for tunnel-wolf effect -8 u Corrected to infinite aspect ratio -8 -4 0 4 8 12 16 20 24 28 32 .4 -2 0 .2 .4 .6 .8 LO L2 /.4 1.6 1.8 Angle of offock, a (degrees) Lift coeffic/ent, 6i F-.. 66.-N.A.C.A. 6406 W.H.

S1a. up}. L'wY.

C C /O O (25 206 -, BB U•C 252.96-L// 40_/0 5.0 4.30 -/./B 5.421 /0 1 -/ .66 O 20 40 60 80 /0 /5 7.78 -.36 Per cenfofchord 20 966 ./7 25 965 69 3010.13 %/2 .44 40 /035 665 SO 9.B/ 486 60 8.78 1.92 2.0 .40 70 7.28 L76 ^p 80 5.34 /.36 U 90 2.95 .74 1.8 .36 95 /.57 35 C, /00 /.091 (-.09) /00 O 1.6 .32 L.E. Rod.: 0.89 v d S/ape ofrodius lhrough end pf 28 p 0 /.4 .28 c chord: sFzo .w /.2V .24 24 20

p. 0

40 . .20 u 420040 A./6 ./6 p /6 60 O O .6 ° .12 £'2 7 80 .4-'.08 8 0 /00 o 4m .2 .04 .o ° l 0 O C 04 Airfoil: NA.CA. 6409 R.N.:3,060,000 c Size: S"x30" 1/eL (f/./sec.): 698 -4 -.2 Pres.lst 3 otm):20.8 Date: 9-/-3/

u

Where tested.• L.0 L. Test: l!D.T 659

-.4 8

Corrected for tunnel-wo//effect -8 -4 O 4 8 /2 16 20 24 28 32 Angle of offock, a (degrees) FILM. 69.-N.A.C.A. 64W .M.11.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 9Yo. up 'r. Uw'r.

C L /0 2 48 /25 C 0 273 -123 U 00.

25 3.60 -1.64 Test No. 739 h o_/0 SO 5.36 -/.99 •// 000 73 657 -,?.

to 7.58 -1.99 0 20 40 60 80 1 00 1S ./0 40 .A18 -1.67 Per cent ofchord -1.25 20 /034 25 11.14 -.76 30 11.65 - .38- 36 v .44 .09 4011.110 .20 s0 /1.16 .55 O 60 9.95 .711 0 .08 32 D 2.0 .40 70 6.23 .65 BO a .73 u 90 333 .39 .07 28 t LB .36 9S /.79 /6 ('12) C /./2) .0 /00 1.6 .32 0 .06 24 p C L.E. Rpd.: 1.511 v l 5 /ooe otrod/us Ci u through end of 1.4 .28c O 28 0 0 .05 20, ^hord:s/zo v v^ /.2.24 : 0 v 24 20 04 16 p ti M v LO 1.206 420E40 .03 12 c y o U

e^./so 16 60 .02

o q .6X./2 .0/ ^ 4 o /2g 80 k .4 .08 8 `0/00 0 0^ k c 4 m 2 .04 C -4 0 ' C .O v _, 2 C ^, _ 8 0` O 0 v 0 R Airfoih N.A.C.A.64/2 R.N:3,090,000 U C Size: S VeG(f//sec.): 69.5 _, 2 -4 4 .3 12 sf7dolm):20.8 Dote:12-22-31 e Airfoi/NA.CA. 6412 R.N.: 3,0917,000 .• Where tested•L.MA.L. Test: V D.T. 739 De:/2-22-31 Test: VD.T 739 _16 E _.4 -B o -.4-- Corrected for tunnel-wall effect 0 1co _ -cfed to infinite os ect rofb -8 -4 0 4 8 /2 /6 20 24 28 32 -4 .2 0 .2 .4 .6 .8 /.O l.2 L4 L6 L8 Angle of attack, a (degrees) Lift coefficienl,C FIX 60. N.A.C.A. 6412airfoil.

$YO. Up'r. L'wr.

D ./2 O - /. w p 10 225 345 -/.53 -1 0 2.5 4.67 - 2.13 4,4 SO 6.44 -2. ]5 p z9 776 -ao6 8.

s /0 0 20 40 60 BO /0 0 o 15 /0.556 -2.97 0 Per cent of chord 20 /L6/ -267 25 12.64 -2.29 W 30 13.15 - L91 44 .09 v 40 1325 -1.25 50 1246 - .76 U V$ BO //. /O - .34 2.0 .40 - .08 70 9./6 -.04

a

80 6.70 .09 r N 90 372 .04 1.8 .36 .u.07 ti 95 2.01 -.05 too 1'161 f-.16/ C III 1.6 .32 0.06 L. E. Rod.: 2.48 C u o Slope ofra0 us through end of 1.4 .28 0.05 28 0 chord: 6/20 W /6 p 24 0 20 1.2V.24^ x.04 C. p.

E - v k o /2'c 20 40 /.0 .0 .20 0.03 L/D y ^ U 8'11 16; 60 .8'.16 0 .02 l o O k l 4s .6,' , .12 .0/ /2U 6 0 k Oo h 8 0/00 .4'.08 0 C ` -4 o o 4 v .2 .04 tj a v o u 'B 0, 0 Q 0 0 ` Q, -. 2 e Airfoil: NA.CA. 6415 R.N:3,060,000 c U Si-5"x30" Vel(ft/sec.): 69.8 -/2 -4 0.

Pres.(sthdotm):208 Dofe:9-3-3/ v ti Where fested.- L.MA.L. Test: V.D.T. 661 Dote:9-3-3/ Test: -8 -.4 0 -.4 Corrected for tunnel-wolf effect Corrected to infinite aspect -4 .2 .4 .6 .8 LO 12 -8 -4 0 4 8 12 16 20 24 28 32 Angle of attack, a (degrees) Lift coefficient,C Frovas 61. - N.A.C.A. 6416 airfoil.

36 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 0 0 ur 4.26 -1.92 / 0 25 1.

25 7.62 -2.59 v.; O 3.0 7.53 -346 p C ].5 9.99 -39/ /0 l0 10.06 -4./5 0 20 40 60 60 /0 05 1202 -4.26 Per cent oFchord 20 /332 -4.0] h 25 14./7 -3.75 14.64 -340 36 m 30 .44 40 14.70 -2.70 l 50 13.60 -205 U^ 60 1224 -447 2.0 .40 32 U 70 /0.1/ -.94 60 7.40 - .54 v u 90 4.12 - 46 .36 .0 28 b k 95 2.24 - .23 100 (./91 (-./9) .0 /00 0 b /.6 .32 0 24 p L E. Rod.. 3.56 C l v U° U Slope ofrodius 1hro 3 qh end of 1.4 .28-, o 20 u 28 0 chord: 6/20 .v a

a

42V.24;^ v 16, 24 0 20 v c.p.

v k 12:c 44 .20 8 q20040 .0 p a `c Bv.16^ o 16 L/D O 0 a ^° o 4r .6X ./2 0 o /2^8 k .4 .68 00 8 v l0 0 a n -4 0 2 .04 C 0 u w C _ 8 0 v 0 0 v ^ 0 a 0 Airfoil: NA.C.A.6418 R.N.:3.100,000 Q Size: 5"x30" Ve%(R../sec.): 69.4 2 - 4 0.

2 c Oate: 9-4-3/ 7: NA.C.A. 6418 R.N: 3.100,000 Pres.(st'nd.ofm.): 20.6 8u Where tesled: L M.A.L. Test: V.O.T 662 9-4-31 Test. VO..T 662 _16 e _4 tied to infinite aspect rcho Corrected for tunnel-wo/t effect .4 .6 .8 10 12 1.4 L6 7.8 -8 -4 0 48 12 16 20 24 28 32 -.4 Lift Coefficienf.0 Angle of attach, a (degrees) FIGURE 62 .-N.A.C.A. 6415.W.11.

Sta. u p " . L' w ' r. ` y o /0 0 - O Ur 0 /.25 5.13 -208 25 6.60 -3.04 m 0 _/O .// 44 5.0 5.65 -4.16 7.5 10.24 -4.81 0 20 40 60 80 t0 /O 1/.52 -5.18 Per cent ofchord 40 /5 /344 -552 20 / 4 .79 -5.49 25 1565 -523 30 16.15 -4.9/ 36 y 44 .09 40 /6.16 -416 50 /5.14 -3.40 U 60 13.44 -2.59 32 D 2.0 .40 08- 70 11.06 -/.93 S 80 8.06 -1.17 v 90 4.51 - .65 28 8 18 .36 .0 .07 k '95 246 - .42 loo (.221 (-.221 .d /00 O U 24 0 AS .32 E3,06 L.E. Rad.: 4.85 u l W Slope of ughradiuI C Thro end of 20 w 1.4 .28 1 0.05 28 p 0 chord: 6/20 W /6 0 .04 24 0 a 0 2 w c. p. k V ^ v v LOy.20 12:c y20 4 0 03 8.c .8x./60 .02 616-1-6 0 L/D , l 01 4 a 4^ .6w .12 .O/ a 12 ^ 8 0 00 tl .4'.08 0 B o t0 C Y -4 0 4. .2 .04 G -. / C 0 .o _8 P k v 0 0 -.2 O 0 Q Q Airfoil: NA.C.A. 642/ R.N.:3,030,000 .0 u Ve1.(fl./sec.): 70.5 14EE: 5"x30" 4 .2 .3 Prss.(sf'7d.olm.):20.4 Ooie:9-4-31 Airfoit: N.A.CA. 6421 U V O. T. 663 Where tested: L.M.A.L.. Test: V.OT. 663 Dole: .9-4 -31 a._ 4 -.4 -8 Corrected far lunnel-wall effect F Corrected to infnits os -.4 -2 0 .2 .4 .6 .8 10 -8 -4 O 48 12 /6 20 24 28 32 Angle of offpch, a (degrees) Lift coefficient. C FIGURE 63. -N.A.C.A. 6421 airfoil.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL s . Up ^. G' Y'.

D /O ./2 1.25 1.36 - .60 C 0 2.5 2.00 - .70 0 _l0 50 301 -.62 7.5 3.65 - .43 4.58 O 20 40 60 0 /0 - . l9 60 / 0 15 5.76 .37 Per centofoh-d ./O 20 6.74 .95 7.49 /.51 30 8.06 2.03 44 .09 36 40 6.66 264 50 8.65 3.35

v

60 8.06 3.48 2.0 .40 Ci.09 70 6.69 321 80 5.17 L52 W 90 2.90 1.43 U /.8 .36.07 95 /.57 73 ?6 ^ 100 (.06) (-.06) /00 - 0 ?4 p /.6 .32 0.06 L.E. Rod.: 0.40 v l V u ,,, ,he ndof 44 28C 0.05 ?0 chord: 6125 C 1,111 0 04 0 20 42, .24^u \y',.04 /6p L D c. p. k v 0120040 1.q.2004.03 U

01660 .d .160

.02 12 80 .6U.12 .0/ 4, 60/0 .4' 4 .08 0 s

C 0 4

u .2 .04 yr .v u m lC ^ 0 Q 0 0 -.2 Airfoi/:N.A.C.A.6506 R.K:,$17Q000 0 C u Size: 5"x30" Ve/. (R./sec.): 68.5 -4 0.

-2 e-3 d. Pres.(stnol-):2/./ Date:4-23-31 -foil: N.A.C.A. 6506 R.N: $170,000 -/2 Where tested: L.MA.L. Test. • V.O.T. 586 te: 4-23-31 Test: V D.T. 586 _/8 -8 v Corrected for tunnel-wall effect rrected to infinite aspect ratio -6 -4 0 4 8 /2 16 20 24 28 32 4 .2 .4 .6 .8 /.0 1.2 /.4 L6 1.8 Angle of attack, a (degrees) Lift coefficient,'C FIOME 64.-N.A.C.A. 8608 airfoil.

c s11. L'w'r.

up'r.

0 - 0 v o /0 t 0 1.25 /.93 - .99 2.5 2.75 - 1 .27 C a 0 50 3.59 -445 7.5 4.97 -/.34 45 O 20 40 60 60 /G O / 5.62 -L /5 7.18 -.96 Per cenl ofchord 20 8.22 - . 9 30 3.56 .5/ .44 4010.11 /.39 50 9.97 2.03 V 60 9.20 2.34 2.0 .40 O 70 7.81 2.30 80 5.85 1.87 W U 90 3.29 /.07 l6 .36 u 95 /.78 53 C /00 (.09) (-.09) / a 0 0 0 16 .32 0 LT Rod.: 0.89 v G ° 5/ape ofrodius 'b /hrough endoi 1.4 .28-, 28 chord: 6 25 N U L2V.24 ^- 24 0 20 /6 p.

y v` /.Ov.20^ o 20 a 40 /2 V u U^ 1t 6 /6° p /8 60 8 0 0 .6u./2 a /2 80 4 u .4'.08 8"0 100 x C .2 .04 C ° 4u o U v 0 0 y OQ 0 Airfoil:NA.C.A.6509 R.N.:3,110.000 .c Size: 5"x30" Ve1.0../sec.): 69.4 -.2 Pres.(51'ld. 11-120.6 Dote: 422-3/ -4, v Where tested.• I'll"L.

Tes1:VD.T.585 Dote: -.4 -8 a Corrected for funnel-wo/l eff ct -8 -4 O 4 8 12 16 20 24 28 32 -.4 - Angle of oNock, a (degrees) L 1ff coetfiaenf, c^ Fmvaa 86 .-N.A.C.A. 8600 airfoil.

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS RLI 1111 5fa LPL-.

a - o tzs 2n -/..l a o a x Test M .// 25 356 -/.82 44 w -/o eoo SO 5.02 -226 y o 75 613 -2.43 O 20 40 60 80 /6 /O 7.06 -2.45 15 857 -2.27 Per cenf ofchord 20 9.69 -1-91

1 25 /650 -1.47

36% 3011.07- .98

.44 .09

4011. 6- .06

d

so 229 7/ U 60 1035 i.21 2.0 .40 G.08 70 876 L39 C BON 6.5 I.24 -°C 90.72 28 e 1.8 .36 .u.07 O .O w Ioo o

24 p

1.6 .32 0.06 c L. E. ROd.: I.SB l W V U 5roae ofrodis Through end of

20 u 1.4 .2B 0.05

28t D

chord: 6/25 • v D

/6 0

24 , 42 0 .2 24 x.04

c.p.

v L/D vk /2.c

^20i4 D 110 4.03

a

o I6g 6 0 .8v.1.6 .02

0 0 k 4y

o B 0 .68 ./2 .0/

u w 0x 0 Bolo 0 .4'.08

a

-4 0 .2 .04

° 4u

C, v

w

-B 0 Q 0 0 -.2

Q Airfoil. • N.A.C.A.6512 R.N.:3,180.000

C size: 5'x30" Vel.(ft./sec.): 68.6

-4 0. -.2 u -.3

P es.(sfndolm):20.9 Dote: /2 23-3/

v

Where lesled: L.M.A.L. Test: V D.T.740 /2-23-31

-8 -.4

Correcied for lunnel-11 effect sled to infinite as -8 -4 0 4 8 /2 16 20 24 28 32 F -4 .4 .6 .8 /.0 Lift coeff/cieni,C 4ng/e of attack, a (degrees) Fiaass 66 .-N.A.C.A. 8613 ailfb 1.

S^ OPr LW r.

l0 U 125 3.26 -1.68 0 2.5 4.40-2.33 4,4 _/0 S. 0 604 -3.04 p Z5 8.34-355 0 20 40 60 80 /L 15 297 -3.56 Per cent ofchord 20 11.14 -3,33 2S12.00 -295 30 /259 -249 .44 40 13.00 -1.52 5012. 2- .62

b

60/450 .0e 2.0 .40 70 9.69 .49 0 80 7.22 .60 Z U 90 405 .36 /.8 .36 Ci 0 /00 (:/5i (-JS) /00 - o a /.6 .32 C L.C. Rod.: 246 N CS D t ough ^dof /.4 .281 28 0 chord: 6/25 v L2 i.24cv o - 24 L.

C c. p.

w v /.0 y.20° N20040 LID, v 16 60 ^ t 0 0 o /2 80 .6 °./2 k 8'.100 .4' .08 a n .2 .04 ° 4u .o 0 0 OQ Airfoi/..NA.CA.6515 R.N:3,100,000 C Size: S"x30" Ve1.(ff./sec.): 69.4 -.2 Pres.(sthd.otm.J:20.11 Dofe: 4-22-3/ v Where lesled•L. .40 Test: V.DT. 583 -.4 -B Correcfed for tunneFwof/effect -6 -4 O 4 8 12 16 20 24 28 32 Angle of attack, a (degrees) Lift coefficienf,C FIOU88 87.-N.A.C.A. 8616 slRoll.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL ./2 w^ /o 48 u °c 0

EFFF

C o -/0 0 20 40 60 60 p er cent ofchord 44 0 / / , 0 , ^ C zss 36 Ls4 Los a 2.0 .40 .41 1 .08 °p 32 05 0

w U

S /.B .36 07 - ° 28 0B1 C v l.6 .32 0.06 m dlls f S 14 .281 0.05 ZO w a I I

24 0

1.2V .24 x.04 c.P. /6 I I I k k m N'0o /.0 4 .20 ^.03 /2 Pf u u,

0 16^

.02 0 8' t ^ 0 4

.6'./2 0/ 4

u k 8 0/ .4'.08 0 0: C

0 4 v

.2 .04 u o l

0 2

0 0 v-.2 Cw-8

Airfoi/: NA.C.A. 65/6 1?.1:3, IOQ000 o

C

Size: 5"x30" _ 2 Ve%(ft./sec.//: 693 -4 R -/2 Pres.(sti]dolm):20.8 D&.%1_31 Airfoi/.•N.A.CA. 65/6 R.N:3,/00,000

w ' -8 u

Where tested.-L.M.A.L. Test:VD.T. 582 _ Dote:4-21-31 Test., V.O. T. 582 4 -16 Corrected for tunnel-wall effect '4 Corrected to infinite aspect ratio 12 16 4 8 20 24 28 32 -4 ..2 0 .2 .4 .6 .8 /0 /.2 /.4 /.6 t.8 ng/e of a//ock, a (degrees) Oft coeff/clent,C

r'. 68 .-N.A.C.A. 6618 airfoiL

5^ . L'^ r. ^ a ^O• UP'n Q 0 /.25 4. ]5 -222 ^1 6.17 2S -3.26 R 0-/0 SO 6.20 -450 75 9. ]0 -524

/o 10.93 -57/ O 20 40 60 80 /6

/512.79 -6./3 Per cent ofchord /0 2014.10 -.6.14 25 3015.60 /505 -59Z N -553 44 .09 36 m -444 4015.63 l 5015. 27 -3.27 601181 -220 2.0 .40 V^..06

32 a

1457 -/.30 70 O 80 6.59 - .6B 90 4.B5 - .32 U

/.8 .36 x.07 28

k 95 2.67 - .23 (22) (:22)

.o

%000 0 /.6 32 0.06 24 p L.E. Rod.: 4.65 C U° u l 6/ope of radius 1 / eo endof z6 /.4 .28 0.05 ED'.

chord: 6/25 v 24 l 21 /.2 V .24 v.04

/68-

c.p.

020041 L0^.20U x.03 R u

/6^ & .02

12 1 Bi .6 U .12

.01 4^

o

L/D u

u

B 0 /0, 4`1 .08

0 00 .2 .04 40

0 4u

.o v

0 Q 0 -.2

0 4

Airfoil: N.A.CA. 6521 R.N.:3,080,000

r U

Size: 5"x30" Vel.(fi!/sec.): 69.8 _4'0. _.2 C_3

Pres.(sfnd.ofm):20.6 Dote: 4-21-31

Airfoi/. N.A.C.A. 6521 R.IV:3,080,000 -8 u Where tested: LM.A.L. Test- VD.T 58/ D ote: 4 2/ 3/Test: VAT 581 Corrected for tunnel-wall effect Corrected to infnite aspect ratio -8 -4 0 4 8 12 16 20 24 28 32 -4 - 0 .2 .4 .6. 8 /.0 12 1.4 16 l Angle of oiiock, a (degrees) Lift coefficient.

Fmm. 68.-N.A.C.A. 6521 M doiL 40 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS L'11 UPr.

^ / S" /.25 2.45 -1.43 M 0 2.5 3.41 -L94 to So 4.78 -247 C o 75 5.64 -270 O 20 40 60 80 /0 /0 6.70 -278 &// -269 Per cent ofchord 20 9.17 -2.41 25 9.95 -200 3010.51 -1.5/ .44 4011.14 -.49 so/ 50//./3 53 0.56 1.44 2.0 .40 70 9.33 L98 BO 7.19 /.90 C 4./6 /./9 90 U /.8 .36 95 2.28 .60 /00 (./L (-.l2/ /00 O 1.6 .32 c L. C. ROd.: L58 N Slope Ifrpdius /h .- end of 1.9 .28 C 28 p 0 chord: 6130 ,d L2V.24 24 L.

Q v LID /.O y.20 20 0 40 /6 60 o rJ .6'./2 o /2.g 80 .4'.06 q,-/00 ix I c C 0 4v .2 .04 o u N 0 0 R 04 R.N.:.1250,000 Airfoil: N.A.C.A. 66/2 C Size: 5"x30" Vel.(N./sec.): 68.0 -4 0.

-.2 Pres. W s 7d.otm.): 21.0 Dote:/2-28-3/ Ci M Where tested.• L.A.L. Test.• D.T. 741 -.4 Corrected for funnel-wall effect -8 -8 -4 0 4 8 l2 16 20 24. 28 32 Angle of ot/ock, a (degrees) Lift coefficient.

Fl°II88 70.-N.A.C.A. 6812.11M.

U

s u ti U .32 C v

V

'28-, 28 p v V .24•^ 24 0 w k v v.20a N 20 u p, v./60 0 16 ^ O ^ O O u /2 o /2 u V

k

.08 8° .04 ° 46 v OQ C -4 0.

U -8 Litt coeff/'aent,C Ang le 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 cD u -2.55 Per cent ofchord 2.62 -296 -3.00 .44 -ass

a

-2.1s 2.0 40 -1.70 s - /.22 u /.8 36 _ k '33 0 -.061 ( D 1.6 .32 o c o. /o v ti /.4 .28`C 28 ^ v 24 0 /.2 .24 c.p.

V k LO• q 20p .20 u w L/D C 'u o• Bv.16o 0/60 ./2 a /2 u .6 0 G 6 0, 4' 08 w 2 .04 ° 4 u o ^ 0 0 0 4 Airfoil: N.A.C.A.0006T R.N.:3,180,000 c -2 Size: 5"k30" uel.(ft./sec.): 68.7 -4 0.

Pres.(sti7d.o/m):20.8 Doie:4-8-31 _ 4 Where fested•L.MA.L. Test:VD.r 554 8v Corrected for tunnel-wall effect 0 4 8 12 /6 20 24 28 32 L ifl coefficient, C, Angle of attack. a (degrees) Ftoo z72. - N.A.O.A. 0006T airfoil.

6fo. Lm, w, ./2 u c° 0 /.25 /.42 - q2 26 1.65 -1.85 ./t S.0 2.30 -2.30

ERR

7.5 2.54 -2.54 /0.

O 20 40 60 80 /0 2.69 -269 .t0 /5 2.87 -28] Per centofchord 20 296 -2.96 -3.00 no 25 3.00 W 30 3.00 -3.00 .44 .09 40 296 -2.96 50 2.79 -2.79 Q 60 2.49-2.49 '.0 .40 1 .08 70 2.04 -204 D 80 /.46 -/.46

W i

90 .77 -.7] .36 ^.07 b !8 95 4/ - .41 0 0 tD0 (.06) (-06) too 0 0 !6 .32 0.06

c

L. E. ROd.: /.19 ^q U N 1.4 .26 0.05 28 ^ !

,v . 3 /6 8 !2V.241^- m.04 24 2i a v /2 LOv.20 0 y 2 i4t U a .lsp .02

.ev

O o M .0/ 4,^ .6 ./2 o 0 0 .4".08 80/0i Cw x k c 2 .04 G-./ -4 0 a ° 4u

.o L v

0 0 -6 a 0 -.2 Oy e .A. 00068 RN.:3.090,000 °u .c " M(H./sec.): 69.5 .3 -4 R tm.):20.8 Dote:4-8-3/ 0006B R.N.:.X6 '2 c Air. foil. N.A.C.A.

d.L.M.A.L. Test: l!D.T.556 Dote: 4 -8 31 Test: Y 6 556 '•4-4 or funnel-wall effect Corrected to infinite aspect ro: -8 u P 0 .2 .4 .6 .8 l.0 12 14 - a - 4 V 4 c4 ca Jc .4 -2 coefficient, Q Angie of oHccA. a (degrees) Lift FlomtE 76.-N.A.O.A. 00068 abfolL REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS sfo.

upr. cwr.

a " 0 0 0 /o 125 -12s as 1.25 L88 -188 l U 0 -/0 5.0 266 -gas 7.5 3.6/ -361 7 /O 4.21 -421 0 20 40 60 80 /L /5 5.09 -509 Per cent olchord 20 5.64 S. -.464 25 -5.92 1 30 6.00 -600 44 .09

40 5.75 -5.75 360

50 5// -5.// 60 4.29-4.29 2.0 .40 <.08 70 339 -339 32 U 80 2.43-, .43 W 90 f.37 -137 u 1.8 .36 95 78 -.78

28 i

/oo 0 w (a2 i

Lao 0 0 0

16 .32 0.06

L.E.ROd: 0.40 v V U

l

/.4 .28 < 0.05 28 0

20 m

• a c.p.

42U.24* -W.04 24 ^ 20

/6p

m '^ 1.01.20$ x.03 V20 40

12^

L/D u u c .8 .16 .02 U 16 60 a 12$ BO .6'.12 .O/ 4^ U 0--.

8 /00

.4.08

x

.2 .04 4 V

-4 0 C .o v

k

0 y

0 0 -.2

Airfoih N.A.CA. 00127 R.N.:3.120,000 .c e U Si- 5"x30" Ve/.(fL/sec.): 69.5 _4 4 P es.(sti+d.otm):20.5 Dole.' 4-3-31 ti

v

Where ies7ed L.M. A.L. Test: V D.T. 548 Dole: 4-3-3/ Test: U. D. T. 5481

-.4 1-.4--- -8

Corrected for tunnel-walleffect Corrected to inf'niie aspect ratio j-

-8 -4 O 4 8 12 16 20 24 28 32 -.4 .2 .2 .4 .6 .8

/.O 1.2 44 L6 1.8 Angle of attach, a (degrees) Lift coefffcient.0 Fioasa 74-N . A.C.A. 0012T SW0.

c l /o 125 ze4 -1 e4 0 - 25 3.69 -3.69 y -10 4.59 50 -4.59 44 R a 75 5.08 S. -5.06 0 40 7 10 S. -536 20 60 80 1C f5 -574 Per cent ofch.rd ./0 20 592 -5.92 25 6.00 -600 30 6.00 -6. 0 .44 .09 40 593 -5.93 CPo 50. 5.59 -559 V$ 60 4.98 -4.96 2.0 .40 - .08 70 4.07 - 4.07 32 80 2.91 -2.91 W 90 1.55 -1.55

95 -83 u /.B .36 .0 .07

63 28 100 0 ( ..12) (-.. /2) loo 0 0 D f.6 .32 0 t06 L.E.Rod.: U v L< 1.4 .28-, .05 o 28 0 20

Do

ma c.p.

/.2V .24* v .04

24 0 20 6

w

v ': 401 .20 0 .03-- 20040 /2 .8x./60 0 /6 D 60 .02 ° 8' 0 0

a /2 80 .0/

u

w w

.4^ .08 8 0 /00 0

C .ro

^. c

.2 .04 4 -4

0 u

v -8'

0 0 m 0 V -.2 Airfoil: N.A.C.A.00128 R.N.:3.060.000 .c oU Size: 5"x30" Ve/.(ft..Isec.)_ 70./ -I2 c Pres.(sthd.otmf:20.3 Date: 4-3/ 4 0.

'3 A/rf0lL• NA.CA. 00/28 R.N.:3.080,000 Where tested: LM.A.L. Tesf: V. D.T.550 -4 Dole: 4-4-3/ _4 Tesf:V. D. T. 550 -8 v Corrected for tunnel-waf/effect _/6 Corrected to Infinite aspect rotio -8 -4 0 4 8 /2 16 20 24 26 32 -.4 :2 0 .2 .4 .6 .8 40 /.2 44 l.8 1.8 Angle of attach, a (degrees) Lifl coefficient,C Rama 76. - N.A.C.A. 00128 airfoil.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL Sro. Up'r L'wr U 25 283 -2.83 -/D SO 4.28 -4.28 p R -.£4/ Z5 5.4/ to 6.32 6.32 0 20 40 60 80 /O - /5 164 -7.64 Per centofchord 8.46 25 B.88 -8.88 30 9.00 -846 -9.00 .44 40 662 -8.62 50 767 -761 U 60 6.44 -6.44 2.0 .40 70 -5.09 S09 0 B0 3.65 - 365 t 90 205 -205 U /.8 .36 95 /./7 -1.17 o l00 (./B) !-.7Bi /o0 0 0 /.6 32 c L. E. Rad.: O.B9 N 1.4 .281 z8 ° 1 C v c. p.

2,j. 04 24021 k v LID /.0120U V 20o4i u a 8 = ./6 v o /6 & ^ t 0 0 .6X./2 a /2 t 8 r U 8,/0 .4 .08 0 4 v .2 .04 C O U 0 0 Airfoi/NA.CA: 00/BT R. N:3,/50,000 C Ve%(ft./sec./: 69.2 Size: S"x30" -.2 Pres.(sfndatin):20.6 Dofe:4-7-3/ 4 0.

u Where tested:L.MA.L. Test:VOT 552 -8 -.4 Corrected for tunnel-watt effect -8 -4 0 4 8 12 /6 20 24 28 32 Ang le of attack. a (degrees) Lift coefficient,C F/evas 78.-N.A.C.A. 081ST slrfolL ^a i^ 30 sod -9. ,44 40 8.89 -6.69 50 6.36 -BAS 60 Z47 -7.47 2.0 .40 - 70 6.11 -6./1 .0 80 4.37 -437 ,N U 90 2.32 -2.32 /6 ,36 U.

4. 95 1.24 /0o (.78) -1.24 o (-.16) 0 0 a /00 /,6 .32 q U L.E.Rad.:]. /5 V C 1.4 O .28% ZB y ap.

L2^.24 24 l 20 40,^ .20 N 20 0 40 L/D P .e Q) 0 /6a 60 o O .6U./2 12. 80 .4-.06 00 6 0 /00- - C a 4 .2 .04 C .o p 4 0 0 y 0 Airfoi/: N.A.CA. 00188 R.N.:3,180,000 1, ' C - 2 0 Size: S'k30 Vel-(ft/sec): 692 Pre y . (sfnd. otm):20.3 Dote: 4 - 6 - u Where tested.•L.M.A.L. Test: V.D.T 551 -6 _ •4 Corrected for tunnel-wall effect 4 6 12 16 20 24 28 32 -8 -4 0 Angle of attack. a (degrees) Frayse 77. - N.A.O.A. 0018B elrfolL REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS , Sto up'r. L.-r _ E /o 2 48 ,. 0 57 i O 2.53./0 -2./7 yy,_/0 5.0 4.29 -286 o 75 5.16 -a28 O 20 40 60 80 / 00 /O 5.84 -3.57 15 6.82 -3.88 Per cent ofchord ./0 40 20 747 -402 25 7.82 -4.06 N 30 -402 .44 .09 36 w 40 776 -3.84 l 50 7.03 -355 5.94 60 -3.16 20 .40 c.06 L"a, 32 70 4.61 -2.72 u BO 316 -2./0 1.63 -4,26 /.8 .36 ll.07 .87 - .74 28 95 21 /00 (.13) i (-.13) rz /00 0 L6 .32 o.06 24 0 L. E. .: Rod56 /. C w Slope of ivdius C through end of 28 0 l.4 .28- .OS 20 m chard : O. /3/ u 24, 20 c.p. 42V.24 i ^ o y.04 /6p " k C L/D q 20Q40 /. 0,§ .20 ^.03 124-c• o /6^ 60 .A.16 .l6 .02 8'^ 12280 .6.120 .0l 4 u LLJ 1 6 _ -/00 .4".08 0 0 $ ° .2 .04 C,H -4 0 ° 4u C o ^ v 0 O w -.2 -g a 0 4 Airfoil: NA.CA.2R/2 R.N.:3,260000 e .c u Size: 5°x30" 3 el.(f1./sec.): 68.3 - 4 -.2 -l2 -•3 0. Pres. (sthd. otm):20.6 Dote:2-26-32 Airfoil.' N.A.C.A.

2R, 12 AR. 3,260, 000 u Where tested.' L.M.A.L. Tess: VD.T. 764 Test: V D. _ _ Dote: 226 32 T. 764 -8 .4 •4 -l6 Corrected for tunnel-wo//effect Corrected to infinite aspect ratio -8 -4 0 4 8 12 /6 20 24 28 32 -4 -2 0 .2 .4 .6 .8 /.0 12 /.4 1.6 1.8 Angle of atlock, a (degrees) Lift coefficient, C Fro- 78.-N.A.C.A. 211 1 I2 airfoil.

Sta. Up'r. L'w'r.

O - Ov k /0 48 JS 230 -/.52 O F O 2.5316 -2.10 4w-/O -276 44 S'' 7.5 5.29 -3.17 ' 10598 -a42 0 20 40 60 00 156.97 -3.74 Per centofchord ./0 40 Z56-3 25 791 -3.97 30800 -400 44 .09 36 40 7.63 -a96 506.73 -.867 44l D 60 S49 -a66 2.0 .40 -.oe 704. -3.27 O a 60- / -2.64 W 90 1.26 -1.63 U 48 .36 ^ .07 95 .66 -.95 OOp 28 10 (.0 (-.13)

100 O - a

1.6 .320 8 ".06 24 0 L.E.itod.: /.SBC v !h / ^ou endof 28 0 /.4 .28 .05---20t chord: O. 153 C.P.

1.2V .24w y.04/6n 24 0 20 o a mk L/D 1.0 1 20 o ° u x.03 20 p 40 .w .

V u o C k a° p /6D 60 .8./60 .OZ 8^

o l2 80

.6 .12 .0/ q^ u 8'-/00 .4'1 .06 0 0 0 = o C .i. 0 4 v .2 .04 vr -. / 3 -4o 0 o u v 0 O "Q:) -.2 _g a 0 Q Airfoil: N.A.CA. 2R2/2 R.N.:3,130,000 o Q c U Size: S"x30" sec.): 69.4 .2 4 4 Ye'TI Pres.(stnd otm.J:20.5 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 Test: V. D. -8 u 0 -.4 765 _/6 -.4 Corrected for tunnel-wol/effect 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.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 45 0 0 0 0 0 o+x NA.C.A. 0012E L251.9 -1.99 /.89-/.89 10 $4 25 261 -261 2.61 -26/ 0 00 0012E 50 .55 -3.55 3.55 -3.55 O 20 40 60 BO 100 20 -4.20 -4.20 4.20 10 I0 /( Per centofcha'd /0469 - 4.69 4.68-4.69 /55.35-5.35 5.35 -535 205.74 -574 5.74-5.74 360) 255.94 -594 594 -594 .44 .D: I .00 -6.00 30.00 -.00 0 40 .74 -55.74 5.74-5.

C L1$ 74 p 4.31-4.31 50 .3/ -4.3/ 2.0 .40 -.01 3z o /.96-/96 60 /.96 -/.96 Q t .50-50 70 .So - .50 W 90.50- .50- .4/ -/.43 /.6 .36 .0; 90 .50 - .50 -3.23 4.30 O - -45-6.5 X8.0' 95 5 .50 5. 7 se i- 04 '4 0 0.01 /.6 .32 98.79 U oe 0011501( -.591 GO ? 0 /4 .28- - ^.0: t

28 0

O E.Rnd. ^.5e C 1

c.p. /. 162

240 20 a.24v ro.0,

^ vk _

m C.P .

/z ^c 20 40 .20 o!o /660 v.16o .8 .0;

o a

a

./2 4,Y

o 12 ^ 80--t- .6

u .08 00 6 0/00 .4 .04 4t

0 C

e v \ 4:_ m 0 8 ^' 0 4

'' C

0 •^ NA 0 .CA.0012 8 0012F R. N: 3,230.000 U "x30" 5Ve%(ft/sec.): 68.1 ./2e -4 0.

Pres.(sti7dotm.):210 Dote:7(V 14-32 `^c

I N. AC 002 a8 0012F, R.N.:3,230,000

Where tested. L.MA.L.

u VD.T. 7519753 Da1e: 1-(9812)-32 Test: I D.T 751 8 753

/6 Corrected for tunnel-wall effect -•' Corrected to infinite ospectmtio -4 0 4 8 /2 /6 20 24 28 32 0 .2 .4 .6 .8 /.0 1.2 /.4 /.6 /.8 ZU Angle of ottock. a (degrees) Lift coefficient, C FIGURE .-N.A.C.A.

80 0012Fo and 0012F, aufoas.

PRECISION largely eliminated during the later tests. For this report, however, the effect of the error from this source A general discussion of the errors and corrections has been minimized by repeating the tests of many of involved in airfoil testing in the variable-density tunnel the airfoils, including all of the symmetrical series is included in reference 8. In connection with this originally reported in reference 2.

report, it was hoped that a more specific discussion of The magnitude of all such accidental errors was the various sources of error and separate estimates of judged from the results of repeat tests of many the various errors might be given. However, after a airfoils, and from the results of approximately 25 careful study of all the measurements it became tests of one airfoil that were made periodically through- apparent that practically all the errors may be regarded out the investigation to check the consistency of the as accidental; that is, of the type the magnitude of measurements. The accidental errors in the results which may best be estimated from the dispersion of the presented in this report are believed to be within the results of independent repeat measurements. The limits indicated in the following table: major portion of these errors is caused by insufficient sensitivity of the balance and manometers, by the a t 0.15o personal error involved in reading mean values of 0.01 slightly fluctuating quantities, and by the error due to CLmax - 0.03 slight surface imperfections in the model. The last is f 0.003 Cm, perhaps the most serious source of error. The models 0.0006 were carefully finished before each test, but the pres- CD1(CL=0) -0.0002 V ence of particles of hard foreign matter in the air stream 0.0015 tended to cause a slight pitting of the leading edge of CD0 (CL -1) - 0.0008 the model during each test. This pitting was probably In addition to the consideration of the accidental the major source of error in connection with the earlier errors, all measurements were carefully analyzed to tests, but it was reduced for the later tests when the consider possible sources of errors of the type that necessity of a more careful inspection of each model would not be apparent from the dispersion of the was appreciated. After a considerable period of results of repeat tests. A rather large (approximately running the particles in the tunnel were found to be- all the air- come lodged, permitting this source of error to be 1.5 percent) error of this type is present in REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS velocity measurements resulting from a reduction in For the purpose of comparing the results from differ- the apparent weight of the manometer liquid when the ent 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 air-stream turbulence. In air streams having different degrees of turbulence, the value of the effects of this error, however, are reduced by the pres- Reynolds Number cannot be considered as a sufficient ence of another error in the air-velocity measurements due to the blocking effects of the model in the tunnel. measure of the effective dynamic scale of the flow The measured coefficients, obtained by dividing the The airfoil characteristics presented in this report were measured forces by %pV 2 , as well as the derived coef- obtained at a value of the Reynolds Number of approx- imately 3,000,000, which corresponds roughly to the ficients are, of course, affected by errors in the air- velocity measurement. Aside from this source of Reynolds Number attained in flight by a medium- error, it is believed that only two other sources need sized airplane flying near its stalling speed. Consid- be considered: first, the deflection of the model and eration of the effects of the turbulence present in the variable-density tunnel (see references 11 and 12) leads, supports under the air load; and second, the inter- ference of the airfoil supports on the airfoil. The however, to the belief that these results are more angle of attack and the moment coefficient are affected nearly directly applicable to the characteristics that by the deflection of the airfoil and supports. The would be obtained in flight at larger values of the Reynolds Number.

error in angle of attack, which is proportional to Cmc l ,, was found to be approximately —0.1° for an N./2 ,ar o ma

t Symmetrii al ai^ foi R ! i I

I 27 per rodlan

x 4% V y + 6% ro

° Maximum thickness in per cent of chord

Camber position in fraction of chord F1aunE 81.—Variation of liftcurve slope with thickness.

(Abscissa of maximum mean-line ordinate) FIGURE 82. — Variation of lifGcurve slope with camber. Results for 12 percent thick airfoil having a moment coefficient of —0.075. The airfoils.

error from this source in the moment coefficient is DISCUSSION inappreciable at zero lift, but at a lift coefficient of 1 may amount to —0.001. The errors resulting from The results of this investigation are here discussed the support interference are more difficult to evaluate, and analyzed to indicate the variation of the aero- but tests of airfoils with different support arrangements dynamic characteristics with variations in thickness lead to the belief that they are within the limits indi- and in mean-line form. For the analysis of the effect cated in the following table: of thickness, test data from consecutive tests of airfoils having different thicknesses and the same mean-line « ±0.05° form are used. The analysis of the effect of the mean- 0.00 CLmax line form is made with respect to consecutive tests of — 0.02 10.001 airfoils of the same thickness (12 percent of the chord) C mcl4 f 0.0002 and related mean-line forms. The results are com- CDa(CL=0) 1 0.0000 pared, where possible, with the results predicted by CDa (CZ =1) ±0.0010 thin-airfoil theory, a summary of which is presented in the appendix.

The tunnel-wall and induced-drag corrections ap- LIFT plied to obtain the airfoil section characteristics might also be treated as sources of systematic errors. Such Lift curve.—In the usual working range of an air- errors need not be considered, however, if the section foil section the lift coefficient may be expressed as a characteristics are defined as the measured character- linear function of the angle of attack istics with certain calculated corrections applied.

o o CL =a (a — aaa) Errors in the tunnel-wall corrections, however, should be considered when the results from different wind where a n is the slope of the lift curve for the wing of tunnels are compared. For consideration of these infinite aspect ratio and a Le is the angle of attack at errors, the reader is referred to references 9 and 10. zero lift.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL The variation of the lift-curve slope with thickness given mean line without altering the camber position.

is shown in figure 81. The points on the figure rep- The theory also predicts an increased negative angle resent the deduced slopes as measured in the angular as the position of the camber moves back along the range of low profile drag. These results confirm chord. The experimental values are compared with previous results (reference 1) in that they show the the theoretical values in figures 83 and 84. The ex- lift-curve slope to decrease with increasing thickness.

The camber has very little effect on the slope, as 003 0 43 o63 indicated in figure 82, although a rearward movement 024 0 44 0 64 0 25 0 45 0 65 of the position of the camber tends to decrease the lot a.9t u, X E v, a v o av `64 v^ g14.& y a, v i Comber position in froction of chord fAbscisso of maximum --line ordinate) 704 8 12 16 20 24 Maximum thickness in percent of chord 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- 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_ ° 'Symmefrfcol series .t 6Z mean 2% mean camber 4%mean camber 0 23 series o 43 series x 24 •• x 44 •• + 25 + 45 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 perimental values lie between 100 and 75 percent, the slope in convenient form for noting the general approximately, of the theoretical values, the depar- trends with respect to variations in thickness and in ture becoming greater with a rearward movement of camber.It will be noted that all values of the slope the position of the camber and with increased thickness lie below the approximate theoretical value for thin (above 9 to 12 percent of the chord). Numerical wings, 2a per radian; the measured values lie between values of the angle of zero lift are given in table III.

95 and 81 percent, approximately, of the theoretical. Maximum lift.—The variation of the maximum lift The angle of zero lift is best analyzed by means of a coefficient with thickness is shown in figure 85. It comparison with that predicted by the theory. Thin- will be noted that the highest values are obtained with moderately thick sections (9 to 12 percent of the chord airfoil theory states that the angle of zero lift is pro- portional to the camber if the camber is varied, as thick, except for the symmetrical sections for which with these related airfoils, by scaling the ordinates of a the highest values are obtained with somewhat thicker s REPORT NATIONAL ADVISORY ;OMMITTEE FOR AERONAUTICS sections). The variation with camber, shown in bar (for airfoils having normal camber positions; figure 86, confirms the expected increase in maximum i.e., 0.3c to 0.5c).

MOMENT lift with camber. The gain is small, however, for the normal positions of the camber, but becomes larger Thin-airfoil theory separates the air forces acting on as the camber moves either rearward or forward. It any airfoil into two parts: First, the forces that pro- will be seen by reference to figure 85 that the camber duce a couple but no lift (they are dependent only on becomes less effective as the thickness is increased.

the shape of the mean line); second, the forces that This reduced effectiveness of the camber is in agree- produce the lift only, the resultant of which acts at ment with a conclusion reached in reference 13 that 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- efficient are given in table IV.

-.04 j.

Air-flow discontinuities.—These and other wind- Theoretico/ curvy ' L5 tunnel tests indicate that at the attitude of maximum 2/ w -.03

^E

U o -.02 -.0/ _L 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 63 024 _1 0 64 0 44 o25 o 45 0 65 o ,.BL E m Qt C• 74 V$ i _L a4 Comber position to rrau/on or enora 4 8 12. 16 20 (Abscissa of maximum mean-line ordinate) Maximum thickness in percent of chord Variation of maximum lift with camber. Results for 12 percent thick FIGURE 88. — FIGURE 88.—Variation of moment at zero lift with thickness. Numbers refer to airfoils. mean-camber designation.

lift the air forces on certain airfoils exhibit sudden a fixed point. We then have in the working range an changes which in many instances result in a serious expression for the total moment taken about any loss of lift. The probable cause of these air-flow dis- point continuities is discussed briefly in reference 13. The Cm = C,ea+nCL stability or instability of the air flow at maximum where Co is the moment coefficient at zero lift and lift may be judged by the character of the lift-curve peaks indicated for the various airfoils. The curves nCL is the additional moment due to lift.

As with the angle of zero lift, the theory states that are classified into three general types as noted in the moment at zero lift is proportional to the camber table IV, but the degree of stability is difficult to and predicts an increase in the magnitude of the judge. It may be generally concluded that improved stability may be obtained by (1) having a small moment as the camber moves back along the chord.

Figures 87 and 88 show the values of the moment leading-edge radius, which causes an early break- coefficient as affected by variations of camber and down of the flow with a consequent low value of the thickness compared with the theoretical values. Re- maximum lift, (2) increasing the thickness (beyond the normal thickness ratios), or (3) increasing the cam- ferring to figure 87, the plotted data indicate that the CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 49 moment coefficients are nearly proportional to the the profile drag as the minim um value plus an addi- camber. It will also be noted that the curves repre- tional drag dependent upon the attitude of the airfoil, senting the ratios of the experimental coefficients to we have in coefficient form the camber are nearly parallel to the equivalent curve

C D = C D,+ ( CDamtn + ACDa)

representing the theoretical ratios except that the curves tend to diverge for positions of the camber The induced-drag coefficient C U, which is computed well back. Figure 88 shows that the experimental by means of the formula given in reference 8, is con- values lie between 87 and 64 percent, approximately, sidered to be independent of the airfoil section. The of the theoretical. Numerical values of the moment variation of the profile-drag coefficient with the shape coefficient at zero lift are given in table V.

variables of the airfoil section is analyzed with respect 0 Ox mean combe .

.04 x 2% a 4% + 6% .02

C &-

Symmetric./ oirfoixe * q

+0.

Norma/cambered oirfoi/s + 4 8 0 16 20 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

in fraction positions (0.3c to 0.6c).

FIGURE 91.— Varlation of minimum profile dmg with thioknom far the symmetrical If the resultant of the lift forces acted exactly airfoils.

through the quarter-chord point, as predicted by the to the variations of the two components of the profile theory of thin airfoils, there would be no additional drag.

moment due to the lift when the moments are taken Minimum profile drag.—The variation of the mini- about this point. The curves of against CL, C ,, mum profile-drag coefficient with thickness for the however, show a slope in the working range which symmetrical sections is shown in figure 91. The cam- indicates that the axis of constant moment is displaced bered sections show the same general variation with somewhat from the quarter-chord point. The factor n thickness but, to avoid-confusion, the results are not represents the amount of this displacement as obtained plotted. The variation of the minimum profile-drag from the deduced slopes of the moment curves in the .00& Mean ca, bar,

Mean camber

V s% .Co a )2 4X 4 % Symmetrical airfoil 2% zr, LL

O .2 .4 s .8

O .2 4 .6 .8 /.O Comber position in fraction of chord

of

Comber position in fraction of chord (Abscissa mu. mean- 1 1ne ordinate)

(Abscissa of maximum mean- line ordinate)

FIGURE 92.—Incmase in minimum profiledmg due to camber. Reseltsforl2pereent where k names 90.—Variation of position of constant moment with camber. Values of n thick airfoils. Values of k for equation Cx,... k+0.0060+0.01 9+01 V, increase fmnquationC..l,—C.,+nCz. Results for 12 percent thick airfoils.

is the in CDs.., due to camber and t is the maximum thickness in fraction of chord.

normal working range. The variation of this dis- coefficient with the profile thickness may be expressed placement with thickness and with camber is shown in by the empirical relation figures 89 and 90. Table VI gives the numerical values. Beyond the stall all the airfoils show a sharp +0.0056+0.01t +0.112

CDamfn =k

increase in the magnitude of the pitching moment.

where t is the thickness ratio and k (which is approxi-

The suddenness of this increase follows the degree of inately constant for sections having the same mean stability' at the stall as indicated by the type of the Line) represents the increase in C,,,.,. above that lift-curve peak.

DRAG computed for the symmetrical section of corresponding thickness. The variation of with camber is The total drag of an airfoil is considered as made up CDanfa as shown in figure 92.

of the induced drag and the profile drag. Considering indicated by the variation of k

50 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS The effect of camber is small except for the highly This function is represented in figure 93 as the curve cambered sections having the maximum camber well determined from the results for the symmetrical air- back. Numerical values of are given in table VII. foils and for the airfoils having a camber of 2 percent CDamr ° of the chord. As the camber is increased, the dis- Additional proflle drag.—The additional profile persion of the plotted points from the curve becomes drag, which is dependent upon the attitude of the air- greater. In general the points above the curve corre- foil, has previously been expressed as a function of the spond to thick sections and sections in which the maxi- lift (reference 4) by the equation mum camber is well back. The departure from the = 0.0062 (CL — C,,.,,)' AC De = C DO — C DOmtn curve becomes greater with increased thickness and where may be called the optimum lift coefficient; CL °n, with a rearward movement of the maximum-camber that is, the lift coefficient corresponding to the mini- position. The points well below the curve correspond mum profile-drag coefficient. This equation holds to the thin airfoils.

.028 .024 x °-1

0 OX mean comber I °

2 x — ., 4 % n .020 6%

J

xX#• 11.012 .008 .004 :° tx .2 .3 .4 .5 .6 .7 .8 .9 40 a-a-, F—E 93.—Addaioml pmdle drag.

Because the additional profile drag is not a simple reasonably well for the normally shaped airfoils at values of the lift coefficient below unity. function of the lift, and also because the results as A convenient practical method of allowing for the presented in figure 93 are difficult to follow, generalized increased values of at moderately high values of curves for the relation CD, the lift coefficient is to include the additional profile OCDe = f (CL— CLavr) drag with the induced drag, as suggested in reference 2. For the symmetrical airfoils of moderate thickness are given in figure 94. These curves are given to the term to be added to the induced-drag coefficient represent more accurately the additional profile drag was given as 0.0062 C L E . The relative importance of for the normally shaped sections.

this term may be better appreciated by considering Optimum lift.—The optimum lift, as defined above, that it represents 11.7 percent of the induced drag of is the value of the lift corresponding to the minimum an elliptical airfoil of aspect ratio 6. The same profile drag. As the determination of this value of the method may also be applied to other airfoils if the lift is largely dependent upon the fairing of the profile- value of the optimum lift is not too large.

drag curves, special curves were faired for this purpose Andrews (reference 14), using the part of these data on enlarged-scale plots corresponding to certain related published in references 2, 4, and 5, suggests for the airfoils grouped together. The values of the optimum additional profile drag the form lift coefficients obtained in this manner are given in table VIII. It may be noted by reference to this table

ODr 1

A CL— C L =f1 D0 that the optimum lift coefficient increases with camber CL _%L, CHARACTERISTICS OF AIRFOIL SECTIONS FRO] I TESTS IN VARIABLE-DENSITY WIND TUNNEL and for the highly cambered sections a definite increase it is not primarily dependent upon the shape of the accompanies a forward movement of the camber.

mean line. Nevertheless it is interesting to compare the optimum lift coefficients with the values included .02 in table VIII representing the theoretical lift coeffi- 2..16.

"eon Max, lhickness 69.

cients at the "ideal" angle of attack for the mean 0 07. comber 4.

.0/ x 29,—••

ma line; i.e., the angle of attack for which the thin-airfoil p o 49, + 6%— theory gives a finite velocity at the nose. (See the u ,*

I. IT

0 appendix.)

" -T1

I 1 1 1 6 410'2-

.02 GENERAL EFFICIENCY The general efficiency of an airfoil cannot be ex- Maximum lhickness 9%

.o/

pressed by means of a single number. The ratio of Symmetrical airfoil 20 `6 4° 2 0° .02 Maximum lhickness /2% k .0/ ^ r E

V g I 0

Meon rnmber 6 4 2 f)- x$.02 ^^ Maximum lhickness /5% G° .0/ ov ^F J B 1 U U a.rA ` . er• O O, .02 i Maximum thickness /B9, o w .0/ $ ^o u .oFmo.

1 4° 2 O•.

.02 Comber powhbg in fraction of chord (Abscissa of maximum mean-tine ordinate/ FIGURE 90.—Variation of CI,_lC o.c. witb camber. Results arc tot 12 pereeut Maximum thickness 219, .Ol thick airfoils.

O the maximum lift to the minimum profile drag is, how- ever, of some value as the measure of the efficiency of O 2 .4 .6 .8 10 /.2 14 an airfoil section. The variation of this ratio with FIGURE 91. — Additional profile dmg as a function of Cy— C.,1.

Results era for sit- thickness is shown in figure 95. The curves of this falls having normal camber positions (0.5c to 0.5c).

figure indicate that the highest values of the ratio are Symmetrical series V^/20 2%mean comber 4%mean comber 6%mean comber v 80 o 43 series o 63 series

0 23 series _j

x 24 x 44 x 64 40 +25 +65 + 4 8 12 16 20 4 6 /2 /6 20 4 8 12 16 20 24 Maximum thickness in percent of chord 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 above 2 percent of the chord and a rearward move- ment of the camber (for the highly cambered sections) seetlon T series Normal 11 11-11 The tend to decrease the value of CL,aee/CDO,afa.

0.40 0006 0.10 1.19 numerical values of the ratio are given in table IX. 0012 .40 1.58 3.80 0018 .89 3.56 7.15 SUPPLEMENTARY AIRFOILS The aerodynamic characteristics of the modified For the purpose of investigating briefly the effects sections are given in figures 72 to 77. These may be of certain shape variables other than those discussed compared with the characteristics of the normal sec- in the main body of the report, 10 supplementary air- tions given in figures 4, 6, and 8. The maximum lift foils were tested. The airfoil sections were as follows: coefficients of the modified and the normal sections 6 symmetrical sections with modified nose shapes, 2 i in figure 97.

are plotted against the leading-edge rad i sections with reflexed mean lines, and 2 sections simu- It is interesting to note that the leading-edge radius is lating those of a wing having a flexible trailing edge.

very critical in its effect on the maximum lift when the radius is small. This critical effect is also indicated Maximum thickness 0.12c by the rapid increase in the maximum lift with increas- ing thickness for the thin sections as shown in figure 85.

/.4 Airfoils with reflexed mean lines.-Previous inves- ^Moximum thickness 0.18c tigations have shown that the pitching moment of cambered airfoils can be reduced by altering the form /.2 of the mean line toward the trailing edge, with a con- sequent loss of maximum lift but only a small reduction /.0 in drag. In order to compare the characteristics of sections of this type with those of the related sections of normal form, two airfoils were developed with the Moximum thickness 006e basic thickness distribution of the N.A.C.A. 0012 dis- posed about certain mean lines of the form given in .6 reference 15 y^=hx(1-x) (1-Tx) .4 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 .2 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. Charac- 2 3 4 5 6 7 O / i-Hnn-eons radus in Der cent of chord teristic curves for the two airfoils are given in figures 78 and 79. The principal characteristics of the sec- F-RE 97.-Variation of maximum lift with nose radius.

tions may be conveniently compared with those of the Airfoils with modified nose shapes.-The airfoils of related symmetrical section, the N.A.C.A. 0012, and the first supplementary group investigated were de- a related normal section having a camber of 2 percent veloped from three of the symmetrical N.A.C.A. family of the chord, the N.A.C.A. 2412, by means of the airfoils: The N.A.C.A. 0006, the N.A.C.A. 0012, and following table arranged in the order of increasing the N.A.C.A. 0018. For each of these basic (or nor- pitching-moment coefficients.

mal) sections one thinner-nosed section, denoted by c.

the suffix T, and one blunter-nosed section, denoted section Cam„ CDC- C.p CDC- by the suffix B, were developed and tested. The derivation of each modified section was similar to that 171 0.004 211M 1.47 0.0086 184 -.002 of the normal section and was accomplished by a sys- 0012 1.53 .0083 2R^12 1.63 .0083 184 -.020 2912 1.02 .0085 IN -.044 tematic change in the equation that defines the normal section. This change is principally a change in the These results indicate that airfoils having reflexed nose radius, but it also results in modifications to the mean lines may be of questionable value because of the profile throughout its length, except at the maximum ordinate and at the trailing edge. The nose radii of the adverse effect of this mean-line shape on the maximum lift coefficient.

sections in percent of the chord are as follows: CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 53 Thickness and camber modifications near the 2. The angle of zero lift moves toward zero with trailing edge.—Two airfoils were developed to simu- increased thickness (above 9 to 12 percent of the chord thickness ratios).

late an airfoil having a flexible trailing edge in a straight and in a given deflected position. The thickness dis- 3. The highest values of the maximum lift are ob- tribution is composed of three parts: the forward por- tained with sections of normal thickness ratios (9 to 15 percent).

tion (0 to 0.3c) having the same distribution as the N.A.C.A. 0012, the rear portion (from 0.7c to the 4. The greatest instability of the air flow at maxi- mum lift is encountered with the moderately thick, trailing edge) having a thin, uniform value, and the central portion joining these two with fair curves. low-cambered sections.

As shown in figure 80, the two airfoils differ only in 5. The magnitude of the moment at zero lift de- creases with increased thickness, varying from 87 to the rear portion, the section designated N.A.C.A.

64 percent, approximately (for normally shaped air- 0012Fo simulating that of a wing having the trailing foils), of the values obtained by thin-airfoil theory.

edge deformed for the high-speed condition, and the section designated N.A.C.A. 0012F I simulating that 6. The axis of constant moment usually passes slightly forward of the quarter-chord point, the dis- of the same wing with the trailing edge bent down in a placement increasing with increased thickness.

circular arc. Curves of the aerodynamic characteris- tics for both conditions are compared in figure 80. 7. The minimum profile drag varies with thickness Considering the results given by both airfoils as two approximately in accordance with the expression conditions for one airfoil, a very high maximum lift =k+0.0056 +0.0lt+0.lt2 C- i.,n with a reasonably low minimum drag is obtained.

k depends upon'the camber and t is

where the value of

CL'-

On this assumption the ratio 197, slightly the ratio of the maximum thickness to the chord.

C."ef. Is 8. The optimum lift coefficient (the lift coefficient higher than the value of this ratio given by the corresponding to the minim um profile-drag coefficient) N.A.C.A. 2412.

approaches zero as the thickness is increased.

In order to study the effects of an extreme change in 9. The ratio of the maximum lift to the minimum the thickness distribution, the principal characteris- profile drag is highest for airfoils of medium thickness tics of the two sections may be compared with those ratios (9 to 12 percent).

of the related normal sections, the N.A.C.A. 0012 and Variation with camber: the N.A.C.A. 6712. The maximum lift coefficient is 1. The slope of the lift curve in the normal working little affected by the change in the thickness distribu- range is little affected by the camber; a slight decrease tion, but it is of interest to note (table I) that the slope in the slope is indicated as the position of the camber of the lift curve of the N.A.C.A. 0012Fa is slightly moves back.

greater than 2a per radian, as compared with an appre- 2. The angle of zero lift is between 100 and 75 per- ciably lower slope for the N.A.C.A. 0012. The profile cent, approximately, of the value given by thin-air- drag is also affected by the change in the thickness foil theory, the smaller departures being for airfoils distribution. Of the two symmetrical sections, the with the normal camber positions.

profile drag of the N.A.C.A. 0012Fp is much higher 3. The maximum lift increases with increased cam- than that of the N.A.C.A. 0012 over the entire lift ber, the increase being more rapid as the camber range. This is not true, however, for the two cam- moves forward or back from a point near the 0.3c bered sections. Comparing the characteristics of the position.

N.A.C.A. 0012F, with those of the N.A.C.A. 6712, we 4. Greater stability of the air flow at maximum lift find that at low values of the lift the profile drag of is obtained with increased camber if the camber is in the former is much higher, but as the lift increases this the normal positions (0.3c to 0.5c).

difference becomes less, and in the high-lift range the 5. The moment at zero lift is nearly proportional to profile drag of the N.A.C.A. 0012F, is considerably the camber. For any given thickness, the difference less than that of the N.A.C.A. 6712.

between the experimental value of the constant of proportionality and the value predicted by thin-airfoil CONCLUSIONS theory is not appreciably affected by the position of The variation of the aerodynamic characteristics of the camber except for the sections having the maximum camber well back, where the difference becomes the related airfoils with the geometric characteristics investigated may be summarized as follows: slightly greater.

Variation with thickness ratio: 6. The axis of constant moment moves forward as 1. The slope of the lift curve in the normal working the camber moves back.

minim profile drag increases with in- range decreases with increased thickness, varying 7. The um creased camber, and also with a rearward movement from 95 to 81 percent, approximately, of the theoretical slope for thin airfoils (2v per radian). of the camber.

54 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 8. The optimum lift cofficient increases with the 4(1-2x) f3(x)- (9) camber and for the highly cambered sections a definite 7[x(1-x)]11 increase accompanies a forward movement of the and y is the ordinate of the mean line at a given abscissa camber.

x. The integrals (4) and (6) may be shown to be 9. The ratio of the maximum lift to the minimum identical with the corresponding integrals given by profile drag tends to decrease with increased camber Glauert (reference 15) and by Munk (reference 17), (above 2 percent of the chord) and with a rearward and integral (5) is given by Theodorsen (reference 16).

movement of the camber (for the highly cambered The evaluation of these integrals for the N.A.C.A.

sections).

airfoil sections given in this report was accomplished analytically. The values of (changed from aL, radians to degrees), C-c4 and CL„ so computed, are given in tables III, V, and VIII, respectively, in the LANGLEY MEMORIAL AERONAUTICAL LABRATORY, main body of the report. This method of evaluation, NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS, however, cannot be applied to many of the commonly LANGLEY FIELD, VA., December 20, 1932.

used sections because they do not have analytically APPENDIX defined mean lines; hence, an approximate method must be used. A graphical determination gives good It is proposed in this section of thereport to present, results and for convenience the values of the three briefly, a summary of the results of the existing thin- functions, (7), (8), and (9), at several values of x, are airfoil theory (based on the section mean line) as ap- given in the following table: plied to the prediction of certain section characteris- tics. Such a summary is desirable because at present x r,(x) f1W f3W x f,(-) W4 W4 the results must be obtained from several different sources which give them in a form not easily applied.

0 0.30 -0.992 0.662 1.111 0.0125 -2.901 118.15 111.17 .40 -1.083 .271 .520 Three characteristics are considered; namely, (1) the .0260 -2.091 39.73 7.747 .60 -1.273 0 0 .0500 -1.537 13.84 5.258 .60 -1.624 -.271 -.620 angle of zero lift aLO, (2) the pitching-moment co- .0750 -1.308 7.403 9.109 .70 -2.315 -.662 -1.111 .1000 -1.178 4.716 3.395 .80 -3.978 -1.492 -1.910 efficient C a,, .15 -1.048 2.447 2.496 .90 -10.61 -4.718 -3.395 d4, and (3) the "ideal" angle of attack .20 2.995 1.492 1.910 .95 -29.21 -13.84 -5.258 or the corresponding lift coefficient .25 -.9so .980 1.470 1.00 that is, values Czl, corresponding to the unique condition for which the theory gives a finite velocity at the nose of the airfoil.

In general, some difficulty would be expected with (See reference 16.)

the graphical method because the values of the above Expressions for lift and moment coefficients may be functions tend to infinity at the leading and trailing written as follows if the angles are measured in radians: edges. Actually, because the ordinates of the mean- line extremities are zero, the integrand may approach Cy=27r(a-%) (1) zero, and does at the leading edge for the integral (4), CLl=27r(aj-a4) and at the leading and trailing edges for the integral (6).

(2) Difficulty, however, is encountered at the trailing edge G 13+aL for the integral (4) and at the leading and trailing edges

me /4 = 2( 0)

(3) for the integral (5). In order to avoid this difficulty, If the leading end of the mean line is chosen as the integral (4) is evaluated graphically from x=0 to origin of coordinates and the trailing end is taken on x=0.95, and the increment contributed by the por- the xaxis at x=1, then the parameters aLo, a,, and tion from x=0.95 to x=1 is determined analytically.

j3 are given by the following integrals Likewise, integral (5) is evaluated graphically from x=0.05 to x=0.95 and analytically for the extremities.

azo = f l y fl (x) dx (4)

The analytical determination of the increments is accomplished by assuming the mean line near the ends to be of the form

ar = f y f2(x) dx (5)

y=a+bx+cc2 t Evaluating the integrals gives yf3(x) dx (6)

d= f

daze = -0.964yo .s5 +0.0954y; (x=0.95 to x=1) where Aa,= +0.467y,.05 +0.0472yo -1 (x=0 to x=0.05) fl(x)= (7) -0.467yo.95 +0.0472y1 (x=0.95 to x=1)

I

7r(1-x) [x(1-x)]E (1- 2x) where yo and yf are the mean-line slopes at the (8) leading and trailing edges, respectively.

f2(x)=2,'[x(1-x)]n CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE -DENSITY WIND TUNNEL REFERENCES 9. Higgins, George J.: The Prediction of Airfoil Characteris- tics. T.R. No. 312, N.A.C.A., 1929.

1. Jacobs, Eastman N., and Anderson, Raymond F.: Large- 10. Knight, Montgomery, and Harris, Thomas A.: Experi- Scale Aerodynamic Characteristics of Airfoils as Tested mental Determination of Jet Boundary,Corrections for in the Variable-Density Wind Tunnel. T.R. No. 352, Airfoil Tests in Four Open Wind Tunnel Jets of Different N.A.C.A., 1930.

Shapes. T.R. No. 361, N.A.C.A., 1930.

2. Jacobs, Eastman N.: Tests of Six Symmetrical Airfoils in 11. Stack, John: Tests in the Variable-Density Wind Tunnel to the Variable-Density Wind Tunnel. T.N. No. 385, Investigate the Effects of Scale and Turbulence on Air- N.A.C.A., 1931.

foil Characteristics. T.N. No. 364, N.A.C.A., 1931.

3. Pinkerton, Robert M.: Effect of Nose Shape on the Char- 12. Dryden, H. L., and Kuethe, A. M.: Effect of Turbulence in acteristics of Symmetrical Airfoils. T.N. No. 386, Wind Tunnel Measurements. T.R. No. 342, N.A.C.A., N.A.C.A., 1931.

1930.

4. Jacobs, Eastman N., and Pinkerton, Robert M.: Tests of 13. Jacobs, Eastman N.: The Aerodynamic Characteristics of N.A.C.A. Airfoils in the Variable-Density Wind Tunnel.

Eight Very Thick Airfoils from Tests in the Variable- Series 43 and 63. T.N. No. 391, N.A.C.A., 1931.

Density Wind Tunnel. T.R. No. 391, N.A.C.A., 1931.

5. Jacobs, Eastman N., and Pinkerton, Robert M.: Tests of 14. Andrews, W. R.: The Estimation of Profile Drag. Flight, N.A.C.A. Airfoils in the Variable-Density Wind Tunnel.

vol. XXIV, no. 25, pp. 530a-530d, 1932 and no. 31, pp.

Series 45 and 65. T.N. No. 392, N.A.C.A., 1931.

710a-710c, 1932.

6. Jacobs, Eastman N., and Pinkerton, Robert M.: Tests of 15. Glauert, H.: The Elements of Aerofoil and Aircrew Theory.

N.A.C.A. Airfoils in the Variable-Density Wind Tunnel.

Cambridge University Press (London), 1926.

Series 44 and 64. T.N. No. 401, N.A.C.A., 1931.

16. Theodorsen, Theodore: On the Theory of Wing Sections 7. Jacobs, Eastman N., and Ward, Kenneth E.: Tests of with Particular Reference to the Lift Distribution.

N.A.C.A. Airfoils in the Variable-Density Wind Tunnel.

T.R. No. 383, N.A.C.A., 1931.

Series 24. T.N. No. 404, N.A.C.A., 1932.

17. Munk, Max M.: Elements of the Wing Section Theory and S. Jacobs, Eastman N., and Abbott, Ira H.: The N.A.C.A.

of the Wing Theorv. T.R. No. 191, N.A.C.A., 1924.

Variable-Density Wind Tunnel. T.R. No.416, N.A.C.A., 1932.

MI i ?a r-.r -.-.8 v eva 2 a 8 t aa cp HIR ...............

sees ................

...........

s 115 tgma tngR ttSS2 22 -8 A I I I I rl I I I I I I I I I I I I - ri I rri I I I I I I rl ^cl - - - - - - - - -9 ...... ; .....................

..... . - . . - .2 - 9 2 2 S 2 -S 2 -S S! 2 2 S -S !2 -2-2.2 -2 e 2 -S 2 c, - ----- . .... .. . .. ....

4aa3^5tl 2 i^k^t-,st al N 52 Rns4 :- . 2 age t3 ^3 iR iR v t S a v tg u a e 8 a! Sg !2 8; ii i2vi & N :,.^ tw v v 4 2:^ & 4 k v Ni i v 4 8 z as 2vM222 t.2 2 .

H M999 9 2i 9 !-f E i R CHARACTERISTIC S OF AIRFOIL SECTIONS FROM TEST$ IN VARIABLE - DENSITY WIND TUNNEL 57 dC^ TABLE IV.-MAXIMUM LIFT COEFFICIENT, CL..

(PER DEG.)

TABLE IL-SLOPE OF LIFT CURVE, ae= dao Thickness designa-

d \ Th' C b d__ - - 13 M A , Z 12

gg 06 12 15 Is 21 25 112 b or des.

tgnntinn 00 .---..-.._._ • 0.98 x 1.27 -1.53 1 1.53 x 1.49 - Laii 1 1.2() -1.53 0.101 0.101 0.100 0.098 0.094 0.089 0.101 011 -___________ 0.102 22 ___ _________ ------- ------- ------- - ----- - - --- ------- ------- 11.60 23 ..-______.__ • 1.04 -1.61 11.60 °1.54 ------- ------- - 22 .103 11.51 1.59 '1: 24 _____.._---- • 1.01 1 1 1.55 1 1.43 -1.36 -::::-: 1 2 - ------- ------- 23:::::::::::: --.-l0,3-- -".-1,0,2-- 7 - 10 - 2 - -------- .101 1 :lot 25 ----------- • 1.03 -1.38 1 1.60 1 1.53 • 1.48 1 1.38 ------- 11.02 :f03 1 l :097 101 26 - ----------- ------- -------------- ------- -------------- - 25 03 02 02 099 ON ------ 02 om.

27------ - --- ------- ------- ------- -------- ------ ------- - :::::: 1: 8 :168 26 ____________ ------- ------- -- ------- ------- -- 00 27 ____________ 42 ____________ ------- ------- ------ ------- ------- ------- ------- • 1.71 43 _______..._. -I : M 160 b1 : 63 lf : 51 11-41 1 : 1 :l 11 1 1 : 42- ----------- ------- ------- ------- ------- - ------- -- 44 11 I 60 ^1.61 17 • 1.47 . 1 17 1.4 43---- ------- :::-, 'OO'D : 45 . .

:11 •1 1.56 1.69 62 A 4 • 6 ------ 1 1.69 ............ .... -- ----- 102 183 10% : 1 , 21 'm : 0 , 03 , . 1,3` 0 44 ------------ : : 46 1171 ---- .. .....

46 ____________ .104 .103 .101 .096 .095 :017 - 47 ____________ • 1. 82 0 ------- - --- 098 47 ____________ ----- -------------- ------- ----- - ------- ------- .097 62_ ___________ ------- ------- ------- ------- ------- ------- ------- 1 1.75 b4 1 1 : 11.87 1l.64 ^l.65 1: 43 1 1.37 11: 11 :, , ::::::: 82.....-_____ -------- - -- ----- -- 64 11 85 : 1 - 41 <I.

- --- - ------- 1 .104 H6 :1 . 7 ig .096 :10I ___._ _ l04 .101 .102 OD9 .099 : OD6 : 101 43 65 ___.__ -1.29 1 1: 6 78 1 •1: 75 16 -1:89 7 -1.61 ^ •1.49 ------- 6 1 1.75 66 --- ------- ------- ------- ------- ------- ------- ----- -- - - 1.83 ....... .

.101 .103 .101 . 099 .095 094 101 OK 1.67------------ --- - -- - -- - --- 097 I Additional tests to determine variation with camber.

TABLE III.-ANGLE OF ZERO LIFT, .4 (DEGREES) Thickness desire- \ t on 06 09 12 15 18 21 25 12 Theor.

b e ll desX ig-tion 00 ___-__._.___ -0.1 0.0 0.0 0.0 0.0 -0.1 0.0 0.0 0 - go 22---- - - - -- - ----- - ----- ---- ----- ---- --------- ---- ' -1. 23. __________- -1.8 -2 . 0 - 9 24. __________- -1.7 -1.7 -1.7 -1.7 -1.9 -1.7 ------ - 1.8 - 2 08 25_ -____.___-_ -2.0 -2 . 0 -20 -1.8 ------ 2 . 1 -2.29 -20 -2.0 - 26 ----------- ------- ------- - 2. 59 NOTE-I,ettff indicates type of lift enrva, peak.

27 -3.04 ------------ ------- ------- - ::::: :::-:: :::::: :.. -N 4 2 ------- ------- ------ ------ ------ ------ ------ 3 .4 -3.60 - ------------ 43 -3.8 -3.6 -3.7 -3.6 -3.5 -3.6 ------ 3 -3.84 8 -3. 9 4. -3. - 8.1 1. -1 _,: 9 -3. 4 --- .15 46......... . _. 3 4.1 - 4. -4.1 -3. 4 ... -4.2 -4.58 40_ ___________ ------- ------- - -4.6 -5.18 -5.0 -6.09 47 ____________ ------- ------- - 92 ----------- ------- ------- ------ ------ ------ ---------- -5.40 - 6 .2 - b .2 -5.4 -5.4 -5.4 - - 8.2 -6.2 ---- -5.5 -5.76 14 -1: 6 -1.1 _ -6.1 _ -^I: 7 -1-7 -I 2 -1-7 -6.23 66...._._.. _. -8.3 6. 3 6. 0 -5.7 -b.3 2 -6.88 66.__________ ------- ------- -- 6 -7.78 -- :::- :::-:: 87_ ___________ ------- ------- 7.0 -9.13 I Based on straight portion of lift curve extended. Bee curve for actual value.

58 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS TABLID V.-MOMENT COEFFICIENT AT ZERO LIFT, C., Thickness 06 09 12 16 18 21 25 12 Theor.

ber des ignation -0.002 -0. ON -0.002 . 002 00_. ._.____. 0.000 -0 - 0.001 -0.003 -0.002 0 2 2 ------------ ------------ ------------ -0.029 -0.0370 ------------ ------------ ------------ ------------ ------------ 23 .__.._.____. -.036 -.038 01 3 6 -.034 ------------ ----- - ------ ----- - --- - -- -.038 - . 0447 24 __._-______ -.039 -.044 W -.010 -.037 -.036 -- - --------- -.044 -.0531 25 -- - --- - ---- -.048 062 -.050 -.019 -.047 -.043 ------------ -.054 om 28_ ___________ ____________ ---- - ------- ------------ ------------ ------------ ------------ ------------ -.000 0749 27 ____________ ------------ ------------ ------------ - - - - -------- ------------ ---- - ------- ------------ - . 076 -.0912 41 ------------ ----- ------ ------------ --- - -------- ------- - ---- ------- - ---- ------------ ___________ _ -.059 -.0739 43 .___.__.____ -.075 -.073 -.072 -.008 -.066 -.057 -------- - --- -.075 -.0994 44 ------------ -.087 - 086 - . 087 -.083 -:078 -.071 ---- : --- : --- -.039 -.1062 46 _-._.__-___ - 109 -:106 -:102 -.097 - . 094 -:082 ---- --- - - - -:105 -.1257 46 ____________ ____________ -.124 -.1497 ------------ ------------ ------------ ------------ ------------ -------- --- 47 ____________ -.143 -.1825 ------------ --- - -------- ---- 62 __ _ - -- - ---- ----- - ------ ------------ -------------------- - --- _________ ------------ ----- - ------ ------- - ---- ------------------------ _.087 -.1109 ------- - ---- --------- - -- 63__ _ __ .__-__. III 110 I il -:106 -:097 -.090 ------------ _.II0 -.1342 64 . . _ __ _.___._ 29 as 29 -.125 -.118 -.110 .... ------- -.132 -.1594 65_ ______ __._ 159 .158 -.154 x-.160 -.139 -.129 __ - . 169 86_.________ :186 2W 67 ----- - ---------- -.206 2737 portion A Based on straight of moment curve attended. See curve far actual value.

TABLE VI.-DISPLACEMENT OF CONSTANT MOMENT TABLE VIII.-OPTIMUM LIFT COEFFICIENT, C,., POSITION IN PERCENT CHORD AHEAD OF QUAR- TER-CHORD POINT (100 TIMES VALUES OF n FOR EQUATION C,+nCL) Cm 04 = Td. Thk__ design s an th"' 00 09 12 15 IS 21 25 12 Cam- Thick- Thickn ess go am bar \d^ Coc a ion i gnatim designs- if.. is Ig • ber des- 00 _____.______ 0.00 0.00 0.00 Q. 00 0.00 0 . 00 0 . 0D0.00 - fgnation 22_ ___________ -- ---- ------- ------ ------ ------ ------ ---- - .17 0.308 23- :: _ ::::::: :17 : 1171 15 272 a - ed 00 _--_.____-_- 0 . 7 0.7 0.9 - 1.1 1.4 IT - 0.9 - ...... . 24 21 1, :1 0 ----- ------ ------ : : :1 20 250 If 18 .16 , , :11 : , ::::: 25-- . IS 08 03 - .23 .251 -- ---- -- 23:::::::::::: ------ -- ------ -- ------ 5-- 3 3 3 20 ____________ ------- ------- ------ ------ --- -- ------ ------ .20 .256 24_ .1 ___________ .3 .4 .7 1.0 1.5 ....... .0 27 ____________ ------- ------- ------ ------ --- -- ------ ------ .20 .272 25 .0 .-_______ ._ .2 .3 .6 1.0 1.7 ------- .7 26 ----------- ------- ------- ------- ------- ------- .8 42 ____________ ------- ---- _ _ ---- ------ --- -- ------ ------ .36 .616 27 ----- ----- ------- ----- - ------- ------- ------- .8 93 ____________ .35 .30 .23 .12 .2 0 .05 ------ .34 .544 -- : 1 :36 : 33 : 22 : 16 : 10 _ _: - : 1 42 ------ ---- ------- ------- ------- ------- ------ to 34 27 22 16 08 30 :50 43_ ____ _______ .3 .4 .5 .7 1.2 1.6 -_________ ------- ------- ------ ------ ------- ------ ------ .30 40 _ .512 44_____ .3 .3 _______ 5 10 L4 1,7 :5 . 1.0 ---------_ ------- ------- -- --- ------ ----------- ------ .33 47 .51A :: -------- -1 .3 : - 8 : 9 1.4 1.7 46_ _____ _____ ----- ------ - 62 ____________ -- ---- ------- ------ ------ ------ ------ ------ .55 .923 ' 47 ___ _ __ _____ ------- ----- I: 63 63 40 M 24 13 47 81 62 ------ ------- . 2 :4 66 6 42 33 21 :117 45 -- - - 6460 : _^:: ^15 -- ------ . 10 - - ----- j 4 1 j 66____ .60 .53 .42 .33 1 M " _gg 64_ ___________ 7 .0 .6 .9 1.3 1.7 ------- .8 86 ___________ _ .0 .0 .7 1.6 667:::: ---- 1.8 1.9 ------- 1.5 6 816 so :' 66__ ___ __ _ ---- 67:-:.::.. 1 ^^ -------- --------------- -------- 1:::::::Il of 2: Theoretical lift coefficient at "Ideal" ougle of attack.

I TABLE V TABLE IX.-RATIO OF MAXIMUM LIFT II.-MINIMUM PROFILE-DRAG COEFFI- COEFFICIENT CIENT, Cna_ TO MINIMUM PROFILE-DRAG COEFFICIENT, CL... 147DO.- Thickness Thick ness _ des na d gs gg ig S.

ig '1" 06 0 9 12 15 18 21 25 12 is i 18 26 12 C bar des, Ignation, I \^d 0.

._.____.--__ 0.. 0074 0.0093 0.0093 0.0108 0.0120 0.0143 ^ 0083 00 -_-____...__ 185 172 184 164 138 115 84 184 ----------- ------- ------- ------- ------- --- --- ------- ------- •0087 22- ----------- ------- ------- -- 23 :0073 : 0083 : 00ag : 0100 ------- - ----- ------- 184 .0089 2a ------------ 142 182 24:. 0070 0080 0090 0099 .0112 .0127 ------ .0095 24 144 26 ___________ .0073 .0081 .0099 .0103 .0112 .0126 _____.___ 189 177 156 128 106 ------- 100 26 14 1 70 180 148 132 109 ------- 184 _ - _ __ 00" 27___ 20 ___ _ ________ ---- ----- - _ ----- - _ ----- ------- ------- ------- 187 _________ 0090 27 ____________ ------- ------- ------- ------- ------- ------- ------- 187 42 ___________ ------- ------- -- ---- ------- ---- -- ------- ------- _ . 0092 42 ____________ ------- ------- - ----- ---- -- ------- ------- ------- 186 49 .____..__.__ .0090 .0089 .0101 .0107 .0119 .0134 ------- .0096 43 --__-____ ___ 161 140 123 44. : - ------- .0076 .0086 .0095 .0105 .0110 .0132 ... __ .0092 96 ------- 172 44 ______-- :::: 1 46 -____.____ . 0037 0003 .0103 .0113 .0125 .0188 ------- .0995 IN ,82 170 150 127 104 ------- 179 46 ____________ ------- ------- ----- - __- ------- ------- ------- 46. .______.___ 132 169 164 143 123 106 ------- 178 _ 0098 47_ ___________ ....... ....... ....... ....... ....... ...... ----- _ 46 ------------ ------- -------------- ------- --------------------- 178 .0104 4> ____________ ------- ------- ------- ------- ------- ------- ------- 175 62_ ___________ ------- ------- ------- ------- ------- ------- ------- - 0101 63__ ------- .0092 .0101 .0108 .0120 .0130 0144 ------- .0102 82.------- - --- 64 __ __i6__ iii _' iiF ii6 ii _ ____ ___ 63- -- -iii :01 :1 :1 :0120 .0132 .0146 ------- .0104 64 .-___--_._-. 160 179 159 Is l a 114 65: 0093 0100 0111 0127 .0141 .0154 of.

--------- ____- - 7 : ----- 65 lag 171 158 132 'a I- 1 ` I. ..__.....__. 114 97 A.

Overview

Document

REPORT No. 460