REPORT No. 460

THE CHARACTERISTICS OF

78 RELATED AIRFOIL SECTIONS FROM TESTS

IN THE VARIABLE-DENSITY WIND TUNNEL

By EASTMAN N. JACOBS, KENNETH E. WARDand 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 theselentificstudy of the problems of flight. Its membership was increased to 15 by act approved March 2, 1929 (Public, No. 908, 70th Congress). It eonsistaof members who are appointed by the President, all of whom serve se, such without compensation.)

JOSEPH S. AMEs, Ph.D., Chairman,President, Johns Hopkins University, Baltimore, Md.

DAVID W. TAYLOR, D. Eng., Vice Chairman,Washington, D.C.

CHARLES G. ABBOT, Sc.D.,Secretary, Smithsonian Institution, Washington, D.C.

LYMAN J. BRIGGS, Ph.D.,Director, Bureau of Standards, Washington, D.C.

ARTHUR B. COOK, Captain, United States Navy,Assistant Chief, Bureau of Aeronautics, Navy Department, Washington, D.C.

WILLIAM F. DURAND, Ph.D.,Professor Emeritus of Mechanical Engineering, Stanford University, California.

BENJAMIN D. FOULOIS, Major General, United States Army,Chief of Air Corps, War Department, Washington, D.C.

HARRY F. GUGGENHEIM, M.A.,Port Washington, Long Island, New York.

ERNEST J. KING, Rear Admiral, United States Navy,Chief, Bureau of Aeronautics, Navy Department, Washington, D.C.

CHARLES A. LINDBERGH, LL.D.,New York City.

WILLIAM P. MACCRACKEN, Jr., Ph.B.,Washington, D.C.

CHARLES F. MARVIN, Sc.D.,Chief, United States Weather Bureau, Washington, D.C.

HENRY C. PRATT, Brigadier General, United States Army.Chief, Mat€riel Division, Air Corps, Wright Field, Dayton, Ohio.

EDWARD P. WARNER, M.S.,Editor "Aviation," New York City.

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

GEORGE W. LEWIs, Director of Aeronautical Research.JOHN F. VICTORY, Secretary.

HENRY J. E. REID, Engineer in Charge, Langley Memorial Aeronautical Laboratory, Langley Field, Va.JOHN J. IDE, Technical Assistant in Europe, Paris, Franed.

EXECUTIVE COMMITTEE

JOSEPH S. AMEs, Chairman.DAVID W. TAYLOR, Vice Chairman.

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

THE CHARACTERISTICS OF 78 RELATED AIRFOIL SECTIONS FROM TESTS IN THEVARIABLE-DENSITY RIND TUNNEL

By EASTMAN N. JACOBS, KENNETH E. WARD, and ROBERT M. PINKERTON

REPRINT OF REPORT No. 460, ORIGINALLY PUBLISHED NOVEMBER 1939

SUMMARY

An investigation of a large group of related airfoilswas made in the N.A.C.A. variable-density wind tunnelat a large value of the Reynolds Number. The tests weremade to provide data that may be directly employed for arational choice of the most suitable airfoil section for agiven application. The variation of the aerodynamiccharacteristics with variations in thickness and mean-lineform were therefore systematically studied.

The related airfoil profiles for this investigation weredeveloped by combining certain profile thickness forms,obtained by varying the maximum thickness of a basicdistribution, with certain mean lines, obtained by varyingthe length and the position of the maximum mean-lineordinate. A number of values of these shape variableswere used to derive a family of airfoils. For the purposesof this investigation the construction and tests were limitedto 68 airfoils of this family. In addition to these, severalsupplementary airfoils have been, included in order tostudy the effects of certain other changes in the form of themean line and in the thickness distribution.

The results are presented in the standard ,graphic formrepresenting the airfoil characteristics for infinite aspectratio and for aspect ratio 6. A table is also given bymeans of which the important characteristics of all theairfoils may be conveniently compared. The variation ofthe aerodynamic characteristics with changes in shape isshown by additional curves and tables. A comparisonis made, where possible, with thin-airfoil theory, asummary of which is presented in an appendix.

INTRODUCTION

The forms of the airfoil sections that are in commonuse today are, directly or indirectly, the result ofinvestigations made at Gottingen of a large number ofairfoils. Previously, airfoils such as the R.A.F. 15and the U.S.A. 27, developed from airfoil profilesinvestigated in England, were widely used. All theseinvestigations, however, were made at low values ofthe Reynolds Number; therefore, the airfoils developedmay not be the optimum ones for full-scale application.More recently a number of airfoils have been tested inthe variable-density wind tunnel at values of theReynolds Number approaching those of flight (refer-

ence 1) but, with the exception of the M-series and aseries of propeller sections, the airfoils have not beensystematically derived in such a way that the resultscould be satisfactorily correlated.

The design of an efficient airplane entails the carefulbalancing of many conflicting requirements. Thisstatement is particularly true of the choice of the wing.Without a knowledge of the variations of the aerody-namic characteristics of the airfoil sections with thevariations of shape that affect the weight of the struc-ture, the designer cannot reach a satisfactory balancebetween the many conflicting requirements.

The purpose of the investigation reported herein wasto obtain the characteristics at a large value of theReynolds Number of a wide variety of related airfoils.The benefits of such a systematic investigation areevident. The results will greatly facilitate the choiceof the most satisfactory airfoil for a given applicationand should eliminate much routine airfoil testing.Finally, because the results may be correlated toindicate the trends of the aerodynamic characteristicswith changes of shape, they may point the way to thedesign of new shapes having better characteristics.

Airfoil profiles may be considered as made up of cer-tain profile-thickness forms disposed about certainmean lines. The major shape variables then becometwo, the thickness form and the mean-line form. Thethickness form is of particular importance from astructural standpoint. On the other hand, the form ofthe mean line determines almost independently someof the most important aerodynamic properties of theairfoil section, e.g., the angle of zero lift and thepitching-moment characteristics.

The related airfoil profiles for this investigation werederived by changing systematically these shape vari-ables. The symmetrical profiles were defined in termsof a basic thickness variation, symmetrical airfoils ofvarying thickness being obtained by the applicationof factors to the basic ordinates. The cambered pro-files were then developed by combining these thicknessforms with various mean lines. The mean lines wereobtained by varying the camber and by varying theshape of the mean line to alter the position of themaximum mean-line ordinate. The maximum ordinate

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

of the mean line 2s referred to throughout this report as thecamber of the airfoil and the position of the maximumordinate of the mean line as the position of the camber.An airfoil, produced as described above, is designated bya number of four digits: the first indicates the camber inpercent of the chord; the second, the position of the camberin tenths of the chord from the leading edge; and the lasttwo, the maximum thickness in percent of the chord.

Thus the N.A.C.A. 2315 airfoil has a maximum camberof 2 percent of the chord at a position 0.3 of the chordfrom the leading edge, and a maximum thickness of 15percent of the chord; the N.A.C.A. 0012 airfoil is asymmetrical airfoil having a maximum thickness of 12percent of the chord.

In addition to the systematic series of airfoils,several supplementary airfoils have been included inorder to study the effects of a few changes in the formof the mean line and in the thickness distribution.

Preliminary results which have been published in-

clude those for 12 symmetrical N.A.C.A. airfoils, the00 series (reference 2) and other sections having differ-ent nose shapes (reference 3); and those for 42 cam-

bered airfoils, the 43 and 63 series (reference 4), the 45and 65 series (reference 5), the 44 and 64 series (refer-

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

If the chord is taken along the x axis from 0 to 1,the ordinates y are given by an equation of the form

f y = ad ^x + ax + a2x2 + a3x' + a4x4

The equation was adjusted to give the desired shapeby imposing the following conditions to determine theconstants:

(1) Maximum ordinate 0.1 at 0.3 chord

x=0.3 y=0.1dy/dx = 0

(2) Ordinate at trailing edge

x=1 y=0.002

(3) Trailing-edge angle

x=1

dy/dx= —0.234

(4) Nose shape

X=0.1 y = 0.078

The following equation satisfying approximately theabove-mentioned conditions represents a profile havinga thickness of approximately 20 percent of the chord.

t y = 0.29690.1/ — 0.12600x — 0.35160x2 + 028430x3— 0.10150x4

. / - ^^olo^° a - N. A. C. A. familyt_e

r

O ^e

+ Clark Yo 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)

5.0Ord____.I O 1 31 -157 4.3581 5.9261 7.0001 10.8051

1&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 windtunnel of the National Advisory Committee for Aero-nautics during the period from April 1931 to February1932.

DESCRIPTION OF AIRFOILS

Well-known airfoils of a certain class including theGottingen 398 and the Clark Y, which have proved tobe efficient, are nearly alike when -their camber isremoved (mean line straightened) and they are reducedto the same maximum thickness. A thickness variationsimilar to that of these airfoils was therefore chosen forthe development of the N.A.C.A. airfoils. An equationdefining the shape was used as a method of producingfair profiles.

This equation was taken to define the basic section.The basic profile and a table of ordinates are given in

figure 1. Points obtained by removing the camberfrom the Gottingen 398 and the Clark Y sections, andapplying a factor to the ordinates of the resultingthickness curves to bring them to the same maximumthickness, are plotted on the above figure for com-parison. Sections having any desired maximum thick-ness were obtained by multiplying the basic ordinates

by the proper factor; that is

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

+0,28430e-0.10150x')

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

where t is the maximum thickness. The leading-edgeradius is found to be

a

r` 2 0 20 ae) =' .Iota

When the mean lines of certain airfoils in commonuse were reduced to the same maximum ordinate andcompared it was found that their shapes were quitedifferent. It was observed, however, that the rangeof shapes could be well covered by assuming somesimple shape and varying the maximum ordinate andits position along the chord. The mean line was,therefore, arbitrarily defined by two parabolic equa-tions of the form

yo= be+ b ix+ b2 X2

where the leading end of the mean line is at the originand the trailing end is on the x axis at x=1. Thevalues of the constants for both equations were then

expressed in terms of the above variables; namely,(1) Mean-line extremities

X=0 y^=0

X=1 y'=0

(2) Maximum ordinate of mean line

x=p (position of maximum ordinate)

and

yr=(1 p)2 [(1-2P)+2Px-el

(aft of maximum ordinate)

The method of combining the thickness forms withthe mean-line forms is best described by means of thediagram in figure 2. The line joining the extremitiesof the mean line is chosen as the chord. Referring tothe diagram, the ordinate yt of the thickness form ismeasured along the perpendicular to the mean line

from a point on the mean line at the station along thechord corresponding to the value of x for which yt

was'computed. The resulting upper and lower surfacepoints are then designated: -

Stations x„ and x,Ordinates y„ and yt

where the subscripts u and l refer to upper and lowersurfaces, respectively. In addition to these symbols,the symbol 0 is employed to designate the angle be-tween the tangent to the mean line and the x axis.This angle is given by

6 = tan-1 dx

y Oa lxa. ya B=Ion-' dx./0- 9 yt

v Mean /insP1

Chordp ^ x

Or /xt, yt/ x„ = x - y, sine Yu ' Y + yt cos B-./OL Rodius fhrough end of chord xt = x + y, sin a yt = yt - yt cos e

/.00

Sample calculations for derivation of N.A.C.A. 6821

z v, ue tan a sin v aos a y, sin e y. cos a z, v. x, m

0 0 0 10.40000 0.37140 0.92840 0 0 ........................ 0 00.01250.3000

0.03314.10503

0.00489.06000

.383330

.357930

.933751

0.011860

0.03094.10503

0.00084.30000.

0.03683.16503

0.01438.30000

-0.02805-.04503

.am1

.07988

.0221.04898

0-.07347-.17143

-.07327-.16897

.99731

.98662-.00585-.0037

.07906

.00218.80585

1.00037.12863.0218

.59415

.99963-.0307-.0218

Slope of radius through and of chord.

FIGURE 2.-Method of calculating ordinates of N.A .C.A. cambered airfoils.

yo=m(maximum ordinate)dye/dx = 0

The resulting equations defining the mean line thenbecame

Myr = p2 I2Px - el

(forward of maximum ordinate)

The following formulas for calculating the ordinates

may now be derived from the diagram:

xa =x-y, sin 6ya =y,+y, cos 0x l =x+yt sin 0y t =yr -y, cos 0

Sample calculations are given in figure 2. The centerfor the leading-edge radius is placed on the tangent tothe mean line at the leading edge.

06

-GX09

001Z-

O55

0-018C -^L

0025

2206 2306 2406

2209 2309 2409

'— 2212_ 2312 ^2412

2215 X315

2218 2318 2418C_^^ ^_' _.^1 '^

-2421

^-----2,506 2606 2706

2509 2609 2709

2612 -^2712

2615 -^

8' ^ 2618 2918

-^

^4206 4306 4406 4506 4606 4706

4209 4309 4409 4509 4609 4709

9212 4312 4412 4512 4612 4712

4215 4315 4415 4515 4615 4715

4218 4318 4418 4518 4618 4718

4221 4321 4421 4521 4621 9721

6206— 6306 6406 6506 6606 6706

6209 6309 6409 6509 6609 6709

6212 6312 6412 6512 6612 6712

6218 6318 6418 6518 6618 ^-6718

6221 6321 6421 6521 6621 6721

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

A family of related airfoils was derived in the mannerdescribed. Seven values of the maximum thickness,0.06, 0.09, 0.12, 0.15, 0.18, 0.21, and 0.25; four valuesof the camber, 0.00, 0.02, 0.04, and 0.06; and six valuesof the position of the camber, 0.2, 0.3, 0.4, 0.5, 0.6, and0.7 were used to derive the related sections of this

family. The profiles of the airfoils derived are shown

collectively in figure 3.For the purposes of this investigation the construc-

tion and tests were limited to 68 of the airfoils. Tablesof ordinates at the standard stations are given in the

figures presenting the aerodynamic characteristics.These ordinates were obtained graphically from thecomputed ordinates for all but the symmetrical sec-

models, which are made of duralumin, have a chordof 5 inches and a span of 30 inches. They were con-structed from the computed ordinates by the methoddescribed in reference 8.

Routine measurements of lift, drag, and pitchingmoment about a point on the chord one quarter of thechord behind its forward end were made at a ReynoldsNumber of approximately 3,000,000 (tank pressure,approximately 20 atmospheres). Groups of airfoilswere first tested to study the variations with thickness,each group containing airfoils of different thicknesses

but having the same mean line. Finally, all airfoilshaving a thickness of 12 percent of the chord were

F—E 3.—N.A.C.A. sWoll pmffiw.

tions. Two sets of trailing-edge ordinates are given.Those inclosed by parentheses, which are given tofacilitate construction, represent ordinates to which

the surfaces are faired. In the construction of the

models the trailing edges were rounded off.Three groups of supplementary airfoils were also

constructed and tested. The derivation of these air-

foils will be considered later with the discussion.

APPARATUS AND, METHODS

A description of the variable-density wind tunneland the method of testing is given in reference 8. The

tested to study the variations with changes in the

mean line.RESULTS

The results are presented in the standard graphic

form (figs. 4 to 80) as coefficients corrected after the

method of reference 8 to give airfoil characteristics forinfinite aspect ratio and aspect ratio 6. Where morethan one test has been used for the analysis, the infiniteaspect ratio characteristics from the earlier test havebeen indicated by additional points on the figure. TableI gives the important characteristics of all the.,airfoils.

Lw0r.

.947 U i 00/2

- /.307/.777

y W -/0y p

-241-2.673-2.669

O 20 40 60 80 100Per cent ofch-d •l0

.44 .09

20 .40 G.OB,y

1.8 .36 u .07.

L6 .32 V

8.06

L4 .26% 0.05.v

/.2V .24 ^ v .04K

40 20 °.03u

.8 8 .16 0 .02o O

60./2 .0/

.4^ .0B 0

.2 .04

0 0 ^-.z

2..00/

-2.902a 7-2.282,832

- /.3/2- .724.4031-0.40

L D c. p.

C

Airfoil: N.A.C.A. 0006 R.N.:.^21Q000Sze: 5"x30" Ve%(0../sec.): 66.5Pres.(sthd.. otm.):20.8 Dofe: l-4-32

_' 2 . 3

Where felted.L.MA.L. Test.• U.D.T. 744Corrected for tunnel-woll effect. -.4 -.4

..40 4 8 1216 20 24 28 32Angle of .,tack, a (degrees)F-- 4.-N.A.C.A. 0000.W.H.

u

01

28 0

24 0k

N 20^

0 16 yl lo /2^

g0

0 4uvoQ.0

-4 W

-9 0Airfoil: N.A.C.A. 0006 R.N-: iDate.-I-4-32 Test: V LCorrected to infinite aspect ro

2 .4 .6 8 10 1.2 1.4Lift coefficient, C

y36 v

l32

28 ^.O

24 020 w

16,

12CIE8 ^^

4,^u

0 0-4 0

m

-8 a

-/2

-/6

sta.O12525

5.07.51015202540503.607060901.

304.50/

Op'r.O1.420,96/666

3./50.5/24.009

4.3034.456352-'

3.4232.746/.967

C.'r.0

-1.-1.96/961-2666-3.150-3.512-4.009-4.303

cuU

v a /0U -C O

U4 0 -/0

O 20 40 60 80Per cent ofchord

-4.456-4.5014.352-3.97/-34232.746

066-1.066-/.967

/00L.E.100!095)

0Rod.:

(A 5J00.69

Cc. p.

LD

11 vi I

C

Airfoil: N.A.CA. 0009 R.N.:3,2/0,000Size: 5 -x30" Ve/.(Alsec.):68.5Pres.(stnd.otm.):20.8 bbafe: /-6-32Where tesled:L.M.A.L. Test:VOT 746Corrected for tunnel-wall effect.

D

SU

41

28

1? 1

2421

q 2004/

o /6D 6it ^a /2^Bi

8 0 /0,

° 4u.o

04c

-4 F-8 u

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

.44 .09

20 .40 ^'.. 08N

1.8 .36 .g,.07

1.6 .32 0.06

e U

1.4 .28- 0.05

1.2G .24 v.04

^

/.0 8.208 Q.03u

.88.160 .020

. 680

.12 .0/w

•4`1 .08,0

.2 .04

0 0 y -.2

°u-.2 ^ -.3

-.4 0 -.4-6 -4 0 4 6 12 16 20 24 28 32 -4

Angle of oftack, a (degrees)Fla". 6.-N.A.C.A. 0000 MIMI.

h36 wl32 ag

ea

24 l20 w

/6 pv/2

8'^k

4^

Op

-4 0

"a ceAirfoil.• N.A. C. A. 0009 R. N.: 32/0,000Dote. /-6-32 Test: V.D.T. 746Corrected to inf)iVM aspect ratio

0 .2 .4 .6 .8 10 /.2 14Lift coefficient. Cl

8 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAIITICS

t

20

v o /o ./2

u O

00 ./0-.44 .09

2.0 .40 -08N

/.8 .36 x.07

1.6 .32 0.06C1 U

/.4 .281 0.05,m V

v.04v ^=

1.01.208 .03

A .16 02o O

..6'.12 .0/

.4'.08 0i

.2 .04

0 0u .2u

-.2 +^ -.3

-.4 -.4-8. -4 0 4 8 /2 16 20 24 26 32 ` -4

Angle of oftoch, o: (degrees)

FI°URE 0.-N.A.C.A. 0012 Oit 0.

U0U

`o

280 0

24 0 20w

g20 40

0 /6 60l ^a /2 80

8,0100

0 4v.g ul

C-4 0.

-8 u

Ot'w

av

26 0 024 0 20-0 40

p /66012 80

9 -/00

° 4 uo u

0 Qc

-4 4

-8 u

Sla Up'r.0 0

2.5 26/55.03. 5557.5 4.200/0 4.683/55.34520 573825 5.941306.00240 5.60350 5.294604.563-4.70 366480 2623901.445-95 .807%00 (. ^)

/.25 /.8.94-1.8940

-2615-3555-4.200-4.683-5345-5.738-5.941-6.002-5603-5294

563

oO 20 40 60 60

Per centofchord

-3.664-2623

1.446- .607(-. p 6)

L. E. Rod.: 1.58 C

e.p.

L/

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

./2

0010

.44 .09

$20 .40 r 08

Cv/.B .36 `.07V

1.6 .32 0.060 u

1.4 .26- 0.05.v

42V .24•wy.04

,.0 208 .113

.8 w .16 0 .02a

.6 ou .12 .0/v

.4v .08 .0t

.2 .04 j-./vO 0 y -.2

u-.2 3-,4 0-.4

-6 -4 0 4 6 12 16 20 24 28 32 4Angle of ot/oc/r, of (degrees)

Flo- 7.-N.A.C.A. 0015 etrlall.

Slo Up'r.

0 0[252.3672.53.2685.0 4.4437.5 5.250/05853/5 6.68/207./7225 7.4230

7. 50

40 7.25450 6.61860 5.70470 4.S80BO 3.27990 /.8/09 /.006

l00 (./SB)/ O

L'w'r

0-2367-3268-4.443-5.250-5653-6.681-7./72- 7.427-7.502-7.254-6.618-5.704-4.580-3279-/.8/0-/.008

(-. 15610

a 20

N p /0U l 0

11

0

0 20 40 60 80Per cent ofchord

L. E. ROd.: 248 C

c. .

/L

F

Airfoih .A. 0015 R.N.:3,200,000 Ve/.(ft./sec.): 68.4fm):2/.O Dote:/2-9-31.• LMAL. Test: VD.T. 728

or tunnel-wall effect

.4 .6 8 1.0 /.2 1.4

Lift coefficient. C.

Lift coefficient C

20

/2>a Up'r. L'w'

c U /O

z

00av

26 0

0

24 L. 20a

y20x40

o /6A 60l01280

8 01000 4m.o l

0 Q-,1'Cry

u-8

44 .09

2.0 .40 GC

08

/.8 .36 x.07

46 .32 0.06

1 U1.4 .28c 0.05

/.21.24 x.04M

/.Oy.20^4.03U a

.8.160 .02

.6 u .12^ .0/q-

.4 -4 .08 0

.2 .04

0 03

-.2

-.4 0-.4k-8 -4 0 4 8 12 16 20 24 28 32 -4

AnO/e of oitock, a (degrees)Fionaa 8.-N.A.C.A. 0018 airfofl.

10 264/Z53.9,11",SO 5.332

6.300to/0 7,02420 6.0/8-20 662569/30 9.003

-2.8412 -.?.922

-5.332-6.300-7.024_6.0/66.606

^c Ov ^ _/0C o

0 40 60 BO /0Per cent ofchord

-6.9/2-9.00340 6.507.84/

04160 6845-6.84570 543660 393590Z/ 7295 1.210

(./089)AX

L.E,Had:

-6.705-7.94/-5.496-3935-2.172-/.2/(-.0 9)

3.56 C

c. p.

L/D

C,

Airfoil:N.A.CA.00/8 R.N.:,1150.000Size: S"x30" Ue(.(H./sec.): 68.8Pres.(stnd. atm.): 20.9 Date: 1-6-32Where tesfed:L.M.A.L. Test.• VD.T, 747Corrected for funnel-wolf effect.

Airfoil: N.A.C,A. 00/8Dote: / - 6 -32Corrected to infinite a4

2 .4 .6 .8 /.0

Lift coefficient.0

48

44.

401360)

l32

28 ti,O

24 pl

20 u/6 p

/2 ig

8 l

k4^

Op

-4 0v

-8"'

.44

R. .40

/.8 .36

/.6 .321

14 .28c.v

1.21.24 1vtE/.0y.20°u

us .l6 0o O

.60t: ./2

.4-'.08

.2 ,04

0 0

-2

-.4

-8 -4 0 4 6 /2 16. 20 24 26 32Angle of attack, of (degrees)

FM 9.-N.A.C.A. 0021 al/3olL27077 0-35--2

Lift coefficient,

St. up'r.

l0 33/425 4.57506.22/7.5 7.350l0 6./95/59.354-9.35420 to 125 39730 540/0/550 9.2660 7.96670 6.4/28041090 253395 /.4/2-/.4/21011(.221,1-.2211/ 0

L'wr.

-3.3/4-4.576-6.22/-7.350-6./°5-/0.04/-/0397-/0.503-/0/55-9.265-7.986

C C /Ou s° 0

k _/0y o0 20. 40 b'0 80 t0

Per centofcho rd

-

-6.9/2 -4590-2.533

OL. n Had.: 4.65

Cc.p.

L/D

a

Airfoil: N.A.C.A. 0021 R.N, 3,190.000Size: 5-"x30" t/ei.(H./sec.): 68.6Pres.(stnd.atm.):20.8 Date:1-7-32Where tested'L.M.A.L. Test: VD.T. 748Corrected for tunnel-wall effect.

lOu0OacN

2B 0 0

24 a 20k

V20 0 40

/6t 60l

/2480

60(000 4 u.o

lC Q

C

-8 u

9CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE -DENSITY WIND TUNNEL

5/a.0

255.07.5/O

20253040506070609095

00

252.44

/57.25

Up'r.

-

3.354.625.55-26.27

7747937.977.667.026.074.90352/.931.05(./3)

LWr.

O-/.46-/.96-235

9-3/-3.44-3.74-3.94-4.03-392-3.56'-3.05-2.43-1.74- .97- .ss(-613)

D 20y o !0V t 0lR

p -/0

O 20 40 60 60 11.Per centofchord

DL.E. Rad: 1.565tope of-d-0

L/ c.p

C

2212 R.N:3,220,000Size: 5"k30" V IMIsec.):68.4P e5.(5ti7dolm.):208 Dote:12-2-31Where tested:L.MA.L. Test.-VDT 719Corrected for tunnel-watt effectj

Aif.il.-N.A.C.A.

agOaW

28 0

0

324 ,b 20020040

lp /6^ 60l l

0 12 80

8'-/00k c0 4 vo u.

l

R 0 Q_4 c0.

u-8

io REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

Up'r.

2S 5^75.0 7.4067S 6.750/0 9.756-9.756/5//./36-/1./3620//.953-/1.95325/2.37830 250440 2.0950/1.029-/1.02960 9.507-9.50770 7.633-7.63380 5.46590 30/61.680-100 (.262)0 0

L'wi:

-5497- 7.406-6.750

z D /0

dV1R

C o-/0O 20 40 60 BO /L

Per centofchord-/2.378-/2 504-209

-5465-30/6/.68f-.262

0LC. Rod.: 6.86

c.p. 11 IC. 11

r-'

C

Airfoil: MA.CA. 0025 R.N.:3,230,000Size: 5"x30 •• Ve/(fl/sec.): 68.3Pres. (sfnd ofm.J. • 20.9 Date: 1-8-32Where tested L.MA.L. Test: VD.T. 749Corrected for tunnel-wall effect

lOtU

OUcv

28 0

1

324 0 21kq2004(

16 61

0 /281

B o /Ofti0 4 m.o u

l

R 0 4.0

4 0.

V

-8

.442.0 .40

1.8 .361.6 .32

G°/.4 .281

rL 2,j. 24

v1.0 ,y .20

A.16v./6^0 0

.6 u./2

.4 08

.2 .04

0 0

-.2-.4

8 -4 0 4 8 12 16 20 24 28 32An of attack, a (degrees)

1'..30.-N.A.O.A. 0026 airfoil.Lift coefficient,

44

2.0 .40d

/.8 .361.6 .32 0U1.4 .28 0

w ^

2j .24 v

101 20v^ o

B` .16 0O O

6u./2

4".08

.2 .04

0 0 w0-.2

`c

a

_40

-8 -4 0 4 6 /2 16 20 24 28 32Angle of ottack, a (degrees)

2 48

ll 44

./0 40

.09 36

.08 32

.07 28

.06 24

x .05 20

.04 16

.03 /2

.02 8'

4

0 p

_.2 -8

.3Airfoil.-NA.CA. 2212 R.N:3,220,000

12

-. 4 -- Dofe:/2-2-3/ Test: VD.T. 7/9 Corrected to infinite ospecfrot/o -lB

.4 -2 0 .2 .4 .6 .8 /.0 /.2 /.4 t.6 18Lift coeff/cient,C

FIGME 11.-N.A.C.A. 2212 airfoll.

2 48

/ 44

/0 40

.09 360l

.06 32 D

.07--28 bO

.O6 24 0

.OS 00,N

.04 /6 003 z

w.02-a 118.0/<

0

0 0^U

O-4 0

v

.3 Airfoi/. N.A.CA. 2306 R.N.:3080,'00 -/2_ q9ote:9-24-31 Test:. V. D. T. 680 _/6

Corrected to infinite aspect ^-otio-4 .2 0 .2 .4 .6 .8 10 /.2 1,4 /.6 18

Lift coefficient.0

1

.44V2.0 .40v

/.8 .36 .9

/.6 .32 0C3

1.4 .28 PN O

/, 2 ti .24m '^

/O.w.20$

eo./sa.6^./2v.4 .^ .08

Y

.2 .04 V`

7

5.0 336 -2.0/7.5 4. 9 -224

20 6.06 -2.5225 6.37 -25130 6.50 -25040 632 -2.3950 5.62 -2.1360 5.07 -1.7670 4./1 -1.3680 2.96 - ,9790 /.64 - .5495 .66 - .33

100 (./01 /-./O)loo - 0

St. up'r, L'^r.m^ l0

I25 1.69 -1.16 U t 02.5 2.39 -/.56

C y _/0

to a. 7 -2.36 0 20 40 60. 80 /015 5.54 -250 Per centofchord

5""' oJrod'usThrough end ofchord: 2175

L.E.ROd.:0.69 C

c. .1

L/D

1

Airfoil: N,A.OA. 2309 R.N.:2,970,000Size: 5'x30" l/e%(R/sec.J: 71,0Ires.(st7d atm.J:203 Dote:9-24-3/Where tested.• L.M.A.L. Test. l!O.T.68/Corrected for tunne%wa//effect

-8 -4 0 4 8 /216 20 24 28 32Angle of attack, a (degrees)

Pim. 13,-N.A.C.A. 2909 airfoil.

.44

R. .40

/.8 .36

1.6 .32

V

/.4 .261.v

1.0 y.20u

8w./6o0 0

.6,"./2

.4 ^ .08

.2 .040 0

-.2

-.4

OU0OacW

280 0

24 0 20w

'20 Q 40

$1660

F12880Bolo

o qmo U

0^.0-q 0.

-8

0

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE -DENSITY WIND TUNNEL 11

0U0c

28 0 0

324 a 20

A20`40

g /6 u 60

u /2 u 60

8'-/0

° 4 vv0 1

.c-4 0.-8 u

0

7.5 30/ -/.22

20 4.65 -1.0925 4,91 -1.0430 S.00 -1.00404.86-.9450 4.49 - .8/s0 3.92 - .6570 3.19 - .4860 230 - .3390 /.26-.1995 66 - .13

l00 (.06) (-.06)/00 - 0

5^ up'r. L'^ r. D /0Z25 1.16 - .73 ° g 025 /.70 -.95 ly5.0 2.43 -/./5 ^ 0-/0/0 3.46 -1.22 0 20 40 60 80 /0/5 4.18 -/./6 Per centofchord

C.E. Rod.:0.40Slope afrod'uslhraugh end ofchord: 2115

C. .

LD C

Ca

Airfoil: N.A.C.A. 2306 R.N.:3,080,000Size; 5"X30" Vel(ft./ser,4:69.8Pres.(sfnd.otmJ206 Dofe:3-24-31Where 1-1.1. L,M.A,I, Test: l<D.T. 660-4 Corrected for tunnel-wall effect

0 0 v0

-.2c

-.4F

-6 -4 0 4 6 12 16 20 24 28 32

An of attack, a (degrees)FIGURE 12 .-N.A.C.A. 23N airfoil.

12

0U0

28 0

024 0 20k

Q 20 Q 40

p 16, 60

l2 g 80

6,0/00

4u0 Q

C-4 0.

-B

O

OUv

0' 20o4

2k

l

p DD o° 6

8 0 /000 4vo u

0 4-4

V-8

2B 0

0

/6 6O

/2 0

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

20

wo /o ./2m k _/^ o• x Test A

St..

0/.252.5507.510/520253040506070809095

/00

up,,.-

2.243.//4.3/5./B5.866.89].547.88B.00].777.146.215.023.622001.09

/ i3

L'w'r.

0-1.57-2/6-285-326-352-3.82-3.94-3.99-3.64-3.45-z.92

y

0 20 40 60 80 /LPer cent ofchord

-231-/.63- .9/5Z

/ 03 CL L. Rod.: 1.58

/harddh2%/dr ofSlope ofrod'us

LID

a

Airfoil: N.A.CA.2312 R.N.:,1120,000Size: S"x30" Ve%(N./sec.): 69.2Pres.(sf'nd.otm.):20.9 ofe:12-2-31Where tested.' L.MA.L. Test.• V.D.T. 720Corrected for tunnet-wall effect

' -8 -4 4 8 /2 /6 20 24 28 32 -4Angle of attack, a (degrees)

F-B 19.-N.A.C.A. 2312.W.R.

Sla.o

Up'r.-

Lw^.6-

20

/ov o 2/.252.5

2003B5

-1.,96-2.74

U ,^. 0l U5 0 5.26 -3.667.5/ 0

6.287.06

-4.25-4.66 O 40 60 60 10n

20 8.97 -538

30 9.50 -5.5040 9.22 -5.2950 7.47 -4.7760 7.36 -4.0670 5.95 -3.2260 4.29 -22890 1.36 -/.2695 L30 .72/00 /.161 (-./6)

/5 8.25 -5.13 Per cent ofchord

(

/00 - O all

25 9.36 -5.48

5/ape afrodiu5

24cho(dh^sot

L.E. Rod.: 2.47 C

0 c.

L/D

C

Airfoil: N.A.C.A. 23/5 R.N.:3,06C108CiSize: 5"x30" Ve1.(fl,/sec.): 69.8Pres.(s%'d.oim): 2Q8 Date: 9-25-31Where tested.• L.M.A.L. Test: Vb. T. 683Corrected for tunne/-wo// effect

Airfoil: N.AC, 23/2 R.A.N.:Dote: 12-2 3l Test:Corrected to infinite aspectO .2 .4 .6 .8 1.0 /.2

Lift co efficient,

44 .09

2.0 .40 G .OBC LL EE

/.8 .36u.07i6 t

1.6 .32 p,06V u

1.4 .28-^ p.OS.v. a

42V.24 •^ v.04w

/.0U .20 u ".03c

.8./60 .020 0

.6 u ./2 .0/

.4^ .08 0

.2 .04 Ok -. /k

O 0 w -.20U

Airfoil: N.A.C.A. 2315_ _.4 Dote: 3-25-31.4 'gCorrected to infinite os

-8 -4 0 4 6 /2 /6 20 24 28 32 .4 :2 0 .2 .4 .6 .8 40Angle of allock. a (degrees) Lift coefficient, CF1omz 16.-N.A.O.A, 2315 Wdoll.

7

.44 .09

2.0 .40 .V^-.OB

/.8 .36 ^.07

L6 .32 0.06

VQ U

1.4 .26": ^.05.d

l2V .24 v.04

40 .20 1)^.03° C

.8 w

.16 0 .020 0

.6^ .12 .0/

.4'4 .08 0

.2 .04 GkO 0 u

-.2

-.4 Q -.4

36 vl

32

28 iO

24l

20

/6 p12

uOx°4 0w

e

48

36 vl

32,

28

/6 0v

/2 c

8e

0

u

040

W

•8 ^`e

00

44 .09If

2.0 .40 06

/.8 .36 vx.07

l.6 .32 0.06C1 U

1.4 .26c th

1.2,j.24.04v

/.O,n .2011 12.03u

.8 m .16 0 .020 0

.6.".12 .0/

.4'.08 0s

.2 .04 C-./O 0 y -.2

0u.2 _3r,.rreC.A.-.4-.4

-8 -4 0 4 8 12 16 20 24 28 32 c4 c2 0 .2 .4 .6 .8 /.0 /.2 /.4Angle of ottock, a (degrees) Lift coefficient.FIGURE 16.-N.A.C.A. 2908 MEN].

N01

U

20 v

/6 p

/2

v4u

0

-4 0v

-8

Sto.0

1.25Z.5.0

101520253050708090

10

404.90604.08

upr:-

/.///.572.28

3243.904.374.694.864.60

2.441.351061

3.33

C w1r.0

-0.60-/.04-1.29- 45-1.44-1.37-1.25- /.12- .90- .70- .49- 20- .1 /(-.061

-B, 10

t 04a-/0p

0 20 40 60 80 /

Per cent ofchord

0.4051ope ofnvdi,5through end ofchord: 2120

I I

LID C

Airfoil: N.A.CA. 2406 R.N.:3, 120,000Size: 5"x30" 3e/.(ft./sec.): 69.3Pres. (sfnd. otm):20.7 Dote: 9-8 - 3/Where tesled.• L.M.A.L. Test. V.D.T. 665Corrected for tunnel-woll effect

°U0

N

2B p 0

24 20

y -0Q40

/6D 60

Q12-co 80

8'.100^. c° 48,

0QC

-4 0.

'B U

9-3/110,000D.T. 666

7

48

44

40

44 .09

2.0 .40 x.08C

/8 .36 x.07

L6 .320.06^q U

1,4 .28 0.05N ^

L2,j.24M \.04v

L0.20 ,03--

.8 W .16 .020 0

.6" .12 .0/

.4".08 0

.2 .04 OE-./

0 0 m -. 2°uv

-.4 0 -.4

166

/2 cw

4sU

0-4 00

.6 PQ

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

13

Oi

O

c0

28 p 03

24 0 20k0 20 0 40

/6a 60

/2g 80

8'.100'• C° 4 u

04C

-4'4U

-B

5^.1.252.55.07.5'/O15RC25304050605.704BO9095

/00too

Up'r./.622273.203.874.435255.816./86.38USS.22. 273.101.72.94

(./0)-

LW-123-166-2./5-2.44-2.60-2.77-2.79-2.74-2.62-2.35-2.02-/.63-124- .85- ,47- ,2B(-./0)

o

-,'b l0U g OY v -/0

O 20 40 60 60 /GPer cent ofchord

G.E. Rod: 0.89/Ih103gh end ofchord: 2/20

C

c. .

L

a

Airfoil., NA.C.A. 2409 R.N.:3,1 /0,000Size. 5"x30" Ve).(ft./sec): 69.3Pres.(sfnd.otm.):20.8 Dofe.3-9-31Where tesfed:L.MA.L. Test.-VDT. 666Corrected for tunnel-wall effect

-6 -4 0 4 8 /2 /6 20 24 28 32 -4 :2 0 .4 .6 .8 /.0Angle of ottock, a (degrees) Lift coeff/cient,C

FIGURE 17, -N.A.C.A. 2109 Wrtoil.

7.5 6.06 -4.47

20 8.70 -5.6625 9.17 -5.7030 9.38 -5.6240 9.2S -5.2550 6.57 -4.6760 7.50 -3.9070 6.10 -30580 4.41 -2./590 2.45 - 41795 1.34 - .66

100 (.16) (-.161too - o

S70o. Up'n GOY: ^^ ^0

Z25 2.71 -206 u t 025 3.7/ -286 y k _/05.0 5.07' -3.84 p

10 6.83 -490 O 20 40 60 80 /L15 7.97 -542 Per centofchord

5/ope ofradiuslhprdh2%20 f

L.E.Rod.. 2.48 C

c. .

LD

Airfoil: NA.CA. 2415 R.N.:3,060,000Size: 5"x30" Ve/(fL/sec.):698Pres. (stud. otm.):20.8 Dote:9-10-31Where tested'L.M.A.L. Test: VD.T.668Corrected for tunnel-wolleffect.

5Iv

28 024 0 20

kC 20' 40

g /s F 60

o /2 8

80/00 4yo u

R 04c_4 0.

-8 u

0

0

14

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

So.!252.5507.5101520253040506070609095

gi

w,

2.152994.13

4.965.636.617.267.677.687.60.7246.365.183.752.061.14

31-

-465-227-3.01-346-3.75-4.10-423

V s° 0vk-/Op

0 20 40 60 80 /6Per cent ofchord

-4.22-4./2-3.60-3.34-2.76-2.14-/.50- .82- .46

'0

L.E.Rad.: 450Slope ofradiusYh -d.- end ofchord: 2/20

C

Airfoi/: N.A.C.A 2412 R.N.:.250,000Size: 5"x30" Ve/. (ft./sec.): 66.0Pres.(stnd.otm.): 2/.0 Dote:/2-3-31Where tested.• &A.L. Test: V.O.T. 721Corrected for funnel-wall effect

-8 -4 0 4 8 12 16 20 24 28 32 -4 -2 0 . .2 .4 `6Angle of ottock. a (degrees) Lift cot

F-E 16.-N.A . C.A. 2412 aic[oll.

a

v28 p 0

24 a 20k

V201 40

0 /6t 60l lo /2.2 80

8 0 /00k

° 4u.o

C 04.c

_4 0.

-6 U

/

.44 .09

2.0 .40 U .08d

48 .36 x.07

/.6 .32 o.06

V U1.4 .28- 0.05

•-/.2V.24^ v.04

v^/.Ov.20 12.03

u.8 u .16 .02o O

.60 .12 .0/

.4 v .08 0

.2 .04 V`-.lk0 0 v -.2U

.2 •3 Airfoil:N.A.C.A-4 0-4 Date: 12-3-31

Corrected to

9-10-31cted to4 .6

Lift co

36 ^

32

26 2f0

24 0l?0 w

16 8-/2ck

Lk

4 ^.OL

°

-4 0

e

jeTest: V. 721 16.sct ratio

10 /.2 1.4 16 /.8^t. C

46

44.

40

N36 N

l32

28 tl

/6 0

/2

4^u0^°-4om

e

./I

44 .09

2.0 .40 G .08v

48 .36 x.07

1.6 .32, o.06^q V

1.4 .28-, 0.05

QUO)

v

.8 v .16 0 .OR--o a

.6 0 .12 .0/k

.4 ^ .08 0

.2 .04 j-/0 0 v -.2

°u-.2 -.3c-.4 0-4--

-6 -4 0 4 6 12 16 20 24 28 32 -4 -2 0Angle of ottock, a (degrees)

Fmv .18.-N.A . C.A. 2415 airfoil.

.44

2.0 .40

LB .36/.6 .32

Ci/.4 .28

vL 2V.24

vLO ,v .2000

U

.6x.1600 0

.6'- ./2

.4.08

.2 .04

0 0

-.2

-.4-8 -4 0 4 8 12 16 20 24 28 32

Angle of attack, Of (degrees)FIOOEE 21.-N.A.C.A. 2421 airfoil.

Lift coefficient,C

sto.0

/.252.55.07.5/015120253012

405060708090as/00(.

/00

UP r.-

3.87S2/7.008.299.28

//.59/2/5

.36

9.7919 4

74

/.7622)-

3.18 -1.66

L'wr.0

-2.82-4.02-SS/-6.48-7./8-6 05-6,52-8.67-8.62

-7.31-6.17-0.671304

-1,06(-.22)0

U cy o /0U r 0

4 a -/00 20 40 60 80 /0

Per cent ofchord

L.E.Rad, 4.855/ape trod,, athro-1gh end ofchord: 2120

dec.P.

/D

Airfoil.-NA.CA.242/ R.N.:3,000000Size. 5"x30" VeL (ft./sec.): 70.5Pres. (sthd.Ol 7.) 0.6 Oate:9-//-3/Where tested. 4.MA.L. Test: VDT. 670Corrected for tunnel-wa/l effect.

UltU0av

28 ° 03

24 ,0 20q20 0 40

o /sa so

o 12 yu° 80

8,0100

° 4 u

C 0 4c

-4 0.

ti-8

v

Q.

ari00l0h0

v.c

k

0v

%C KA.0 A. 24189-11-31clad to inhi,it.4 .6 .8

Lift coeff/cie/ C

7

.44 .09

2.0 .40 x'•.08W

/.8 .36 x.07

L6 .32 0.06G U

1.4 .28`c 0.05v

/.2V .24k m.04w^v

40v.208 L..03u

.8v.161 .02--0 0 -

.6./2 .0/k

.4 08 0

.2 .04 Of-./

0 0 w -.20-.2 -.3

-.4 0 -.4-8 -4 0 4 8 12 /6 20 24 28 32 -4

Angle of .Hock. a (degrees)n-RE 20.-N.A.C.A. 2418 airfoil.

48

44

40

36 ml

32 a

28 tf.o

24 0l

20 u4/6 0

/z c.k

8.

4su0^`o

'4 0v

_8 rn

V

Si..L252.5507.5/0/5

506070BO9095

/00/00

20/0./52510.653010.6840/0.71

up'r.3.284.456.037./78.059.34

9.698.657.02SOB28//.55/./9)-

GWr-2.45-3.44-4.68-5.48-6.03-6.74-7.09-7/8-7./2-6.7/-599-5.04-3.97-280-L53- .67(-./9)O

/0u r 0l U0-/O

0 20 40 60 80 /LPer centofchord

L. E. Rod: 3565"p, ofm ' end a

rad'ust,chord: 2/20

0

a.P.

L/D

n

Airfoil. NA.C.A. 2418 R.N.:3,060,000Size: 5'x30" Ve/(f1/sec): 69.8Pres. (sfnd .1-):20.8 Date: 9-11-3/Where tested.-L.MA.L Test: VDT. 669Corrected for tunnel-wall effect.

DO

4'.

v28 0

0

24 0 20k

20 40N ^0 6D60

a 1280

8'.100

0 4v0 Cc

-4 0.

u-8

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 15

U0tu00acv

28 03

24kq 20F

0 /6

o /2v 80

4u.oOQ

C-4 0.

u-6

.40

.36

.32

G.28-,

vv .24.0h

Wv.20^.0 b,v./6o0 0U.1208

.040

16

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

.44

1.57-1.270

224-2.56

2.9e Per -a ofrhord

An of allock, a (degrees)FIGURE 22.-N.A.C.A. 2608 airfoil.

r. L'w'r.N p /O •/2^$ D

O 20 40 60 80 /00B -2.76 . /D

O

vdfi o

0

24 0 20Q

Q 200400 /6^- 60

.2 l2 809 o /00

° 4u0 Q

.c-4

4u

-B

5/0.0252.55.07.5/0/520253040506070809095

100/00G. E.S/opethrohchord:

Up'

Z2

3-'03.75W. 2SOS5.60-3.5.965/B6.275975.354.44326.99

(.10)-

Rod.:

ug

02-2.97-2.84-2A4-/.97-/.50

ofredius

2/25

-/.06- .67 .36

- .22(-./O)

o0.69

e ndofG$.

0

Airfoil:N.A.C.A.2509 R.N.:,R060,000Size: S"x30" 3e%(ft./sec.J: 69.8Tres. (sfnd. otm.J: 20.7 Dote:30"

teSMCd L.M.A. L. Test: U.D.9-1

T. 673Corrected for funnel-wa//effect

44 .09

2.0 .40 t'..OBC

/.B .36 3.07/.6 .32 x.06

/.4 .28 .05/.2V.24u v.04

c/.Ov.20, x.03

u ^.8,) ./6 0 .02.6 U .12^ .0/.4v .08 0.2 .04 Vr -. /

3.2 _3-.4 E _ 4'

__H

36

322824: p

got/60/28'^4U0j

°-4 0

°

Y.

C"°/' ve

-/2_16Dote:

Airfoil:NA.CA.Corrected9-/531to

2509infinite Test:aspect

R.N.3.060,000l!O.T.ratio 673

-6 -4 0 4 8 12 16 20 24 26 32 c4 c2 0 .2 .4 .6 .8 /.0 /.2Angle of oHOCk, a (degrees) Lift coeff/cient,C

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

.44

2.0 .40

L0 .36

46 .32UQ

1.4 .28N

L 2V .24m

LO ,y .20 0

Al° o

.6X.12k

.4'.08

.2 .04

0 0

-.2-.4

-8 -4 0 4 8 12 16 20 24 28 32Angle of attack, a (degrees)

51a.0/.252.55.07.5/01520253040506070809095

/oo/00L.E.S/apelhroC&

Op".-2.633.634.965.9/6.66

7.758.486.929/99./66.627.646.284.572.56L41!./sl

-RW..:of,ugh2/25

0-2.11-2,94-3.96-4.60-5.06-5.63-5.87-5.92-5.84-535-462-3.76-2.88-/.9B-/.07-.62!-.lsl02.41ai-

end of

u

U /0_10C O

0

NI#R20 40 60 80 /LPer cent ofchord

i

c. p.

LD

C

Airfoi/.• N.A.CA.25/5 R.N:3,o60,000Size: 5"x30" Vet. fft/sec.): 69 8Pres.(stndot. ):20.6 Doi-S-18-31Where tested.-L.M.A.L. Test.-VDT 675Corrected for tunnel-wall effect

lOU0OCN

26 0

324 0 2(k

q 2004(l

o 6D&

0/28(k 8 010(

C0 4v.o u

l0 y

C-4 0.

ti-B

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

r0aa,

26p 0

24 R20kq20a 40

l61 60

0 /2 86

8 0 /OO^- c° 48

O pC

-4 0.-8 u

3.59ssz4304.43

i o -/0 ^ll.1LIJ_LJJ_J-L11J0 20 40 60 80 /00

Per cent ofchorr(

. / /

. l0

.44 .03

2.0 .40 • .OB

1.8 .36 6.07k

1.6 .32Ci 8 .06

1,4 .26^ 0.05

1.2V.241 m•040

LO x.200 Q.0.3u a

.0.160 .02Q)o O

.6X ./2 .01

.4'.08 0

.2 .040 0

O .2

_ _' 2 v "?

-• 4 0 -•4-.4

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

o v q 674 44

40y

36 v

U32 $i

28 2(p

24 0l

20

160v

12c

8lo

u00

-4 0vU8 e

_/2

-16

4.44

%so3.29zs3 oa/.9733

:42/31

C

e Afinsof

LID

I

c.a,

0^H.D

Airfoil: N.A.C.A.25 /2 R.N.:3,080,000Size: 5"x30" VeG/flfsec.:69.4Pres.(sfnd.ofm):21.0 Dote: /2-3-3/Where tested: L.htA.L. Te5t. , VD.T. 722Corrected for funnel-woll effect.

Airfoil:Date:Corrected

N.A.CA. 2512 R.N:3080,000/2-3-3/ Test: V. D. T. 722

/o infinite aspect ratio4 8 12 16 20 24 28 32

ng/e of oHock, a (degrees)FrovaE

..2 0 .2 4 .6 8 10 /.2 1.4 /.6 l.8Liff coefficient, C

FiomE 26.-N.A.C.A. 2616 airfoil.

270770-$6--3

Uru0v

28

24 lu

kA 20o

0 /6^

.2 /2 u

Bo

0 4u.o

0 yc-4 0.U

-B

36

32 aaP

28 2f6

24 p

20 m

16 1-11.

/2

k4uU

00

-4 0m

-6",e

-/2

-16

Angle of attack, a (degrees)

1'..27.-N.A.C.A. 2521 WdOU.

10ote: 9-2/ 3/I Corrected to infinite0 .2 .4 .6 .8 1

Lift coefficient

.44 .097 // SV

s 2.0 .40 -.OB

112`L2 B7 5

3

7 LB .36 x.076J

[. Rpd.: 4.es 1.6 .32 x.06JPe Ofrod'3S

f C^j

o^d^'z% s 1.4 .28- 0.05va

/.2V.24 y.04C.P._ k

4.03

.e^U 11

.lso .020 0

.6'./2 .0/

.4v .08 0

CD .2 .04 OE -. /0 0 -.2

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

tested.• LM.A.L.

_ ' 2 .3Where Test: l!O.T. 677 -.4 -.4Corrected for tunnel-wa/l effect ^

-4 0 4 6 /2 /6 20 24 28 32 :4

36 v

32 O

28.6

24 0l

20",v

168-/2Bi

k4^

0

-4 0v

e

ratio

U

av

2824 0k

420

i

60.0 4wux l

04r

-4 0.v-8

18

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

J/

.44 .09S

.40 .08v.36 x.07

.32 0.06V u

.26`c 0.05•v a

Ci.24v N.04v

v .20 $ 4.03u

m /6 p

. 0.02 aoo O .u ./2 .0/k

1 .08 0

.04

0 m -.2uu

m .3 Airfoil. N.A.C.A. 2518

o _-4 Dote: 9-19 3/Corrected to infinite -

__ .4 c2 0 .2 .4 .6 .8 /.0Angle of attack, a (degrees) Litt coefficient,FIGURE 2B.-N.A.C.A. 2618 airfoil.

m o /0 2u ^ 00-/0

0 20 40 60 BO /00Per centafchord ^ ./G

7

.44

2.0 .40

/.8 .36/.6 .32

V/.4 .281

v1.2,j.24 ;O^

vLO x.208

u o.sv.ls^0 0

.6X./2ti.

.4 08

.2 .04

O 0

-.2

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

Angle of oitock, a (degrees)

Slo.0

[252.55.07.510/520253040506070809095/00

-2.052.663.951725.346.256677.267.517.567.236.565.564.152.341.26Ly

0-1.73-2.37-a/7-a66-403-4.45-4.6/-462-4.52-4.03-3.36-2.56-/.77/

-.56- .33

u U /0Yo' 0

1 U4 p -/0

0 20 40 60 80 /CPer cenfofchord

L.E. Rod.: /.56

c'.9 -,ofchoror 2/30

LID c. p.

0

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

al0uUkO

v28

324 1- 2(

kq20o 4(

0 /sD s(

o 12u 8(

8 104

4 C

0 4

h 04C

-q ti

V-8

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

Dl0lU

O

v

26 o

G3

2402(v

y 10 Q 410 /6 6C

8(

60/a0 4m.o u

0 p.c

-4 0.

ti-8

-tl 4

Fxomz 28.-N.A.C.A. 2612 tWoll.

7

.44

2.0 .40

1.8 .36

1.6 .32

Ci

1.4 .28c.W

1.2, .24

/.o'.20

.8 m .160 0

.6 ./2

.4^ .08

.2 .04

0 0

-2

-.40 4 8 M /6 20 24 28 32Angle of attack, a (degrees)

M.. 29.-N.A.C.A. 2712 sftM.

6t.L25L56. 07.510/5202530405060709095100

804.42

UPY.2042623.904.665.266.136.737./27367.437.136.525.572571.44(113)

L'^Y.-17S-240-323-3.76-4.12-456-4.75-4.77-4.67-4.18-3.47-2.6/

v P /0U i Oh 0-10

0 20 40 60 60 /CPer cent of 'd

-1.67- .83- .33- .19(-. 3) O

L.ERod: /.58.l1h ough end.(chord: 2/35

1

LID c. .

Airfoil: N.A.C.A. 2712 R.M-$060.000Size: S"x30" Vel.(ft./sec.): 69.5P es.(sfnd. 0 Dote:12-4-31Where tested.• L.M.A.L. Test: V.D.T. 724Corrected for tunnel-wall effect

Lift coefficient.(

Slo.O

/.252.55.07.5to152025304050

6070809095

/00L.E.5lopethrou

d:chor

LPL,-

3.044.135756.%Z909.159.749.929.949.56875

Z6.134.392.4//.3/1.131

Rod.:

gh end-4//0

L'wi-.

0-/.07-441-L69-/.66-46/-1.55-1.74-496-2.06-2.05-1.85-/.55

cu

v o /0U t 0l Uh

p -/0

0 20 40 60 80 /6Per cent ofchord

ofrodius

-1.21- .BS- .50- .3/(-63)

756

of

Li

III

L/D c. p.

0

Airfoil: N.A.CA.4212 R.N:3240,000Siie: 5"x30" Vel.(f//sec.): 68./Pres.(sYnd.afm): 20.9 Dote:l2-/0-3/Where tel...^L.M.A.L. Tezt:VD.T. 729Corrected for tunnel-wo// effect

Lift coeff%c%eni, C

-.4-6 -4 0 4 8 12 16 20 24 28 32

Angle of attack, o: (degrees)Hausa 30.-N.A.C.A. 4212 Udall.

.44

2.0 .40

48 .36

1.6 .32Ci

/.4 .,?8<,N

2,j. 24 i?v

/0^.20uU

.8m./600 0

.61" ./2

.4 '.08

.2 .04

0 O

-.2

Dt0Uk

v28 i 0

024

a 20kA 20x40

l6 16 60

80UB 0/00

M0 4o u

2 0 4-4.

00.u

-B

44

2.0 .40 VcwL8 .36

1.6 .32 0U° u

1.4 .28, o/. 2V.24^ m

tiv

1.0,t .20 40

O ^.6x .12k.4 .08

.2 .04 C

0 0 v°u

.2cF_•4 0

-8 -4 0 4 8 12 /6 20 24 28 32 S`Angle of attack, a (degrees)

7.5 393 -.35

20 6.44 .6825 6.65 B930 7.00 %00

40 6.62 /.0250 6.33 /.0360 5.56 9970 4.54 .86BO 329 .6590 /. ]9 34s5 .97 /5100 LO61 (-.O6)

/00 O

Sta. Vp'r. L'w'r A 200 - 0 N p

/.25 /,39 -.57 02.5 2.05 -.64 ti -/050 3./O -.54 p ffffH/0 4.63 -.12 0 20 40 60 80 /6/5 S. .32 Per cent of chord

L.E.Rod.:04051-through end ofchord: 4115

0L/D I I I C

c.p.

n

Airfoil. N.A.C.A.4306 R.N-W,90,000Size: 5'X30" VeL (f1./sec.): 70./Pres.(s £nd. of 3 Od1.:4-11-31Where lested.• G.M.A.L. Test: V.O.T. 561Corrected for funnel-wall effect

D0U

DW

'80 0

24,0 2

q20 0 40

o /6 P

60lo 128

80/000

4 U

OQc

-4 0.-8 U

d

20

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

12 48

44

./0 40

09 36

08 n, 32D

.07 28 t

06 14:;z0l

OS 20 m

.04-/6 p

.03 /2'ck

.02 n 8. C

O/ 4 t0

0 00u

------ 4 0w

Q. 3 Airfoil N.A.C.A. 4306 R.N.:3060.000 _/2

-.4 4-11-31 Test: V. D. T. 561'4 Corrected to infhde aspect ratio _/6-.4 -2 0 .2 .4 .6 .6 /.0 /.2 /.4 /.6 1.8

Lift coefficienl,CFluvaa 31.-N.A.C.A. UN aidoll.

D

`oUc

28 024 0

0200t

p /6U^ t

p /2 U

wo qvo

u

a 00.C-4 0.

-8

LlOlUkO

m

26 ;p

24 O 2fw

q20p41

/6' 61t

o /2 ^ 81

60 /Ol

0 4m.o u

v04

C-4 R

-8

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 21

cows

- 34- %J7-1.28- 1.0/

o i00

o -l00 20 40 60 60 /00

Per cenfofchard

./2

.l0

44 09$

2.0 40 1.08w).8 .36 u.07w

1.6 .32 0.06cj U

/.4 .28 c 0.0542V .24 v.04

Q) 04Ouv.90^

0.03

.8 i .16^ 02

.6 u ./2^ .0/k

.4 .08 0

J-..2 .04 Ci - /0 O

u2

-2'2 ^ '3

-4 ff_4..4

32.-N.A.C.A. 4308 Wdoll.

48

44

40

36 poi

32O

28O

24 pt

20 uCL

16p

C8t

4yU01)

-4 0W

-B tz

-/2

_/6

- 60-_ so29

Co,-- 1s- Os

ao

.05-(-.09)

:T8_9 Crodi3s

fs

n

c ^°n

Airfoil: N.A.CA. 4309 R.N.:3,060,000Size: 5'X30" Ve%(H./sec.): 69.8Pres(stnd. otm.): 20.B D&e.4-/3-31Where tested.• L.M. A.L. Test.• VD..T 563Corrected for tunnel-wall effect Dote:

Airfoi/:N.A.CA.Corrected

4-133/to

4309infinite Test:

aspectV.

R.N.:3060,000ra{io

D. T. 563

0 4 8 /2 /6 20 24 26 32Angle of o{tock, a (degrees)

Fioun

-2 0 .2 .4Lift

.6coefficient.

.8 1.0 1.2 /.4 1.6 18

51a.O

L25 2.642.5

5.07.5/0/520253o40506070609095

100

Up'r.-

3.63

S6.227/28469.349.62/0.009.758.987656.394622541.38(' 13)

L'wY.0-1.29

-/. 75-2.19-2.34-2.39-231-2.17-2.07-2.00-1.86-/.6/-1.28- .95- .64- .38- .25(-Q3)

U cN p

/0

U ^ 00 _/o

0 20 40 60 80 /LPer cent fchord

CL. E. Rod.: 1.58si oe orroe^sthrough end ofchord: 4//S

LID C.P.

a

Airfoil: N.A.C.A. 4312 R.N.:3,/80.000Size: 5"x30" Ve/.(fl./sec.: 68.7Pres.(s{i+d.otm):20.8 Dole:12-14-31Where tested: L.MA.L. Test: V D.T. 731Corrected for tune/-wall effect

)

.44

2.0 .40/8 .36/.6 .32

Ci1.4 .28 c

N

/.2V .24v

LOv.20^u0 0

.4^ .08

.2 .04

0 0

-.2-.4

-8 -4 0 4 8 12 /6 20 24 . 28 32Angle of o{lack, a (degrees)FxoII z 33.-N.A.C.A. 4312 aWoll.

/I

.44 .09S

2.0 .40 G.08,v

1.8 .36 x.07

1.6 .32 0.06V u

44 .28`c 00 .05va

1.2 Ci.24X j.04m o

/. 0,E .20 4.03u

.8 v./600 .02o 'o

s °./2 .0/k

.4".08 0

.2 .04 G-./k0 0 0 -.2

-.2 w-.3v

-.4 Z-4-6 -4 0 4 8 12 16 20 24 28 32

Angle of otfock, a (degrees)

48

44

40

36l

3228 i?

.d

24,l

20

v

`c

4^

00

-4ov

-8

Airfoil: N.A.CA. 4315

Lift coefficient, Q

2.5 4.47 -2.267.5 737 -3.3/

20 /0.60 -a6025 11,32-3.55301/. -3.504011.18 -3.3350/0. / -2.9360 9.01 -2.4270 73/ -L66

BO0

529 -/.309 292 - .7495 1.5B - .44

100 (.l6) (-.16)

sto. UPS'. 4M, D t0/.25 3.32 - 60 u 0

5.0 6./3 -2.97 y o

v

v k -/0/0 0.36 -3.50 0 20 40 60 BO t0l5 9.85 -3.62 Per Cent of,or,

100 0 C

L. E. Rad.: 2.48Slope afradiusthrough end ofchord:4 /5

c. p.L/D

0

Airfoil., N.A.CA. 4315 R.1i:3,120,000Size: 5"x30" Vel.(ft./sec.): 69.3Pres.(stnd. aim): 20.8 Dole:4-14-31Where Msled.•) A.L. Test.-VDT 565Corrected for funnel-wall effect

L-c°kDW

28 0

0

24 0 20

420040N ^e /6 60

/2^8B o /0

e4

mz° 0 4C-4 R

V-8

0

0

2512. 80

Sto. Up 'r. L0,

Z58.56-4.25

20 /22530 13.00 -50040 /264 -47750 //.65 -4.2460 10.15 -3.5570 6.24 -2.7660 5.95 -/.94

90 329 -/.OB95 1.79 - .64100 (./9) (-./9)too - oL. C. Nod.: 356Through end ofchord: 4 /5

0 - 0 W b0/.25 4.07 -/.90 t 02.5 S.36-2.74 y -/0S. 7.20 -3.68 0

/0 9.64 -4.56 0 20 40 60 W /0/5 //.22 -4.92 Per centofchord

Slope ofrod/us L;,

c.p.

LD

C

Airfoil: N.A. C.A. 4318 R.N:3,09Q000Si- 5'k30" VeZ(ft/sec.): 69.5Pres.(st^7datm.):20.8 Dofe:4-14-31Where tested.-L.MA.L. 7 -est.-VDT 566Corrected for tunnel-wall effect

D0kOav

28 p 0

24 3 20

q 200040N ^o /6' 60

0 /2 oC 80u8 0100

0 4v.o u

4 04C4 0.

-B u

22

REPORT NATIONAL ADVISORY COMMPPTEE FOR AERONAUTICS

FIOREE 39.-N.A.C.A. 9315 Elrtoll.

)

.44V$

2.0 .40v

48 36

/.6 .32 0Ci u

1.4 .28 ova

1.2V.24w m

40.^.20va 0u a4

0 0

.6' .12

.4'.08.

.2 .04 V`

0 0 yU

-.2v

-4 0

-8 -4 0 4 8 12 16 20 24 28 32Angle of attack, a (degrees) Lift coefficient, C

FtomE 35 -N.A.C.A. 4318 MO.

sf.JS2.55.0i0/520253040so6070809095/0000

Up',1.251.682.79ass5.155.906.426.766.906.555.954853.561.961.05(.06)-

L''wr.-.64- 79- .82- .60-.25

.12

.46

.741./01.241.271.16.49

24(-o06)

D /00

u u 0u,A

O 20 40 60 80Per cent ofchord

/G

L. E. od.: 0.4051ape ofrodi35through end ofchord: 4120

CL/D

c.p.

IinC,

Airfoil.-N.A.CA.4406 R.N:3,10Q00054, 5"x30" Vef.(ft/sec.): 69.4Pres.(sfd.nofm.):20.9 Dote:8-21-31Where tested: L.MA.L. Test: VOT. 651Corrected for tunnel-wa//effect

D

ukOw

2B o (

24a 2(

O 2004(

/6 6(^ t

u

80/a`^ C

° 4 u,o

a^ 0 Qc

-4 4

U-B

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENBFPY WIND TUNNEL 23

D

tkOD

v

28 0 0

24020

20 40N

/6O 60

o /2 °C80uM 8'.100

C0 4w° u

^° O 4C

-4 0.

8 u

Sto:0

1.252550

,0/0.152530405060//.70809095

/00/00

2013.72

4.841296.26

1262/4.291450/4.0912.96/9./86.643.672.00(.221-

o-2./9-3./8-4.38-564-6.19

S. -42-6.50-6.50-6.23-5.56-4.66

ut / 0lR

o -/00 20 40 60 80 /C

Per cen/ofchord

-3.66-258--1.43

.84(-.22)

4.o

L.E. Rod. 65s/ape 1'=,ihrou9chord: 4115 C

c.p.

LID

vi

C

Airfoik NA.CA. 4321 R.N:3/20,000Size. 5"x30" Ve%(f/./sec.): 694Pres.(sfnd.ofm.):20.7 Dote:4-15-WWhere lested.-L.MA.L. Test: V.OT 567Corrected for funnel-wo/1 effect

./2

7

.44 .09a

2.0 .40 V .08N

18 .36k

1.6 .32 0.06C°

1.4 .28-, 0.05W ^

1.2V .24 x.04a '^

/.0- .20 u x.03u a

.8 v '16° .02o O

.6X./2 .0/k

.4".08 0Y

2 .04 G f

0 0 m-.20-.2 -.3c4 1-4

44

h36 m

32 O

28 b

20 w

/6 0a

4y

0O'er0

-4 0

-81

^A/rfo/1.NA.CA 4321 R.N. 3,120, 0001Dole:4-15-3/ Test, VD.7.5671Corrected to infinite ospectrobo

0 .2 4 .6 8 /.0 /2 1.4 16 LBL/ft coeff/'c/enl, C

-8 -4 0 4 8 12 16 20 24 28 32 -4Angle of oltock, a (degrees)

Ftovaa 38.-N.A .C.A. 9321 SIM%

.44 .09a

2.0 .40 U.08N

1.8 .36 ^.07

1.6 ,32 0.06

V1.4 .28- 0.05

W/.2V.24 v.04

mlo-.20u x.03u

.8 ^./6 0 .02

.6 0 ./2 .0/k

.4".08 0

.2 .04 0-./O O LIZ -.2

U

-.4 0-.48 -4 0 4 B 12 16 20 24 28 32 -4

Angle of oltock, a (degrees)F[0vae 37.-N.A.C.A. 4908 airfo.

48

44

40

36 w

32

28 ti.d

246l

20'.

168-a

12'^

4su

0ZO-4ov

e

N.A.CA. 4406 R.N..:3,/00,000-21-31 Test: VO.T. 651

L ift coeff/'c%ni. 0,

0

44

2.0 .40

1.8 .36 .^

1.6 .32 0V u

1.4 .28 c oW a

V ^

LO,v.20uU

.$ v./s o0 0

.6 U.12w

.4'.08

.2 .04 t

0 0 m02

c

-.4-8 -4 0 4 8 12 16 20 24 28 32

S. 0 3.74 - /.6575 4.64 -1.74

20 7.33 - /.3025 7190 -L0230 8.25 - .7640.8.35- .3550 767 - .0760 7.00 .1470 5.76 .2660 4.21 .2690 2.33 .1495 1.26 03

100 (.03) 1-.091/00 0L. E. Rod.: 0.69Slope o(rodiusthrough end ofchord: 4 20

Sto. up-,. LWr. a cu0 - O u o 10

/.z5 /.e/ -LOS 02.5 2.61 -1.37 R o-/0/0 5.37 -/.73 0 20 40 60 BO //5 6.52 -155 Per centofchord

c.p.

Lm

0

C

AirfOlL'NA.CA.4403 R.N:3,170,000Size: 5"x30" Ve1.(ft./sec.): 68.6Pros.(stndatirz): 2/.0 Dote:8-24-31Where tesfed L.

'WA.L. Test: V DT 652

Corrected for tunnel-watt effect

D

0av

28 'b 03

24 a 20

020040y ^p 16 60

o /2.g 8

8 0/00c

° 4 u00 4

c-4 0.

V-6

0

.44

2.0 .40

/.8 .36

1.6 .32V

1.4 .28W

A2V.24v

LO y.20°U u

$ .!60 0

.6'./2

.4'.08

.2 .04

0 0

-.2

-.4-8 -4 0 4 8 12 16 20 24 28 32 -

Angle of attack, or (degrees)1NOME 30.-N.A.C.A. 4412 91doil.

Lift coefficient. 0.

1.25 2.44 -/.43

7.5 5.76 -2.74

20 6.60 -2.7425 9.41 -25030 9.76 -22640 .460 -/.8050 919 -1.4060 8.14 -L0070 669 -.6580 4.89 - .39so 271 - .zz95 /.47 - ./6

100 - OL.E. Rod.: 1.565/ape of-d,11,

0 lhrough endpfchord: 4 20

S1o. Opi-. Lm, a c

o -0 ^r / 02.5 3.39 -1.95 y5.0 4.73 -2.49 h 0 -/0

/0 6.59 -2.66 0 20 40 60 60 /0

/5 7.69 -2.66 Per centofchord

l00 1.131 (-./31 C

LID c.p.

C

Airfoil, N.A.CA. 4412 R.N:3,200,000Size: 5"x30" Vel.(ft./sec): 66.6Pres.(sfnd.oim.):20.8 Doie:12-15-31Where tested: L.h1A.L. Test: V D.T. 732Corrected for tunnel-wail effect

0ukacm

2810324 1- 20w

Q20 a4l

ri 16 6

12 8

80/0`^ C° 4 u.0

R° 0 4C

_4 0.

_g v

0

0

0

0

24 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

33

ai

2 48

44

0 40

09 36

.08--eat 32.

.07 28

.06 24

.05 20

.614 16

.03

.02 r° 8.

0/ 4'

0 0

-4

-.2 -8c

.3 Airfoil. NA.C.A. 4409 R.N.:3,170,000 12_ 4 Dote: 8-24-31 Test: V. D. T. 652 -/6

Corrected to inf'nite aspect ratio-4 -2 0 .2 .4 .6 .8 /.0 /.2 1.4 16 /.B

Angle of attack, a (degrees)

Lift coeff/cteni,C

Fmm. 38.-N.A.C.A. 4408 I& M.

2 48

44

./0 40

.09 36

.08 °° 32.

.07 28

.06 24'

^.0$ 20

04 l6

.03 2

.02 a 8'

.0/ 4'

0 0

-.2 -8

Airfoil. NA.CA . 4415 R.M.: 3,110.000 /2_ Dote:8-26-31 Test: l!D.T. 654

' 4Corrected to in fihite aspect rofo -16-4 720 .2 .4 .6 .B LO L2 /.4 L6 /.8

Lift coeff can't C

75 6.91 -3.7/

20 /0.25 -415 -25 /0.92 -3.9B30 //.25 -3.75

//.25 -3.255 /0.53 -2.7260 9.30 -21470 7.63 -/.5580 5.55 -/.0390 308 - .5795 1.67 - .36

00 (.16) (-.16)

S0. UP r. L'0r. ^D l0/.ZS 3.07 -1.79 U 02.5 4./7 -2.48 l U5.0 5.74 -3.27 4 p

-l0

/0 7.84 -3.98 0 20 40 60 BO /015 9.27 -4./8 Per cent ofchard

L. C. Rad.: 2.48 Lcslope o7u9through end o7Chord: 4120

e.p.ZT

L/D

oe

Airfoil: NA.C.A.44/5 R.N:3 /10.000Si- S"x30" Vel.(fl/sec.): 69.5P(sti7d.otm):20.7 Dote:8-26-31Zr.Whe tested: L.MA.L. Test: VOT. 654

Corrected for tunnel-wall effect _4 0

-8 -4 0 4 8 • 12 16 20 24 28 32Angle of attack, a (degrees)

.44

2.0 .40 U_6

LB .36 ._6

/.6 .32 0V u

1.4 .28 oW ^

/.2 V . 24k m/.0.200

A./6^0 0

.6u .l2

.4'.08

.2 .04 c

0 0 m02 cv

DO

UkO

CW

28 i 03 .24 'a 20

q20040

0 16--60

x 80/0`• C0 4 v.o ua0 0 4

.c44

-8 u

0

0

0 /0

44 .09a

2.0 .40 ^ 08N

AS .36 ^.07

1.6 .32 0.06G u

/.4 .08- 0.05

1.2V .24 •F v .04v `^

40 .y .200 ^.03.0

.8v.16^ 02-0 0

6 u .12 .0/k

.4 08 03

. 2 .04

0 0 m -.20-.2 u-.3

v-.4 0 -.4

-8 -4 0 4 8 12 16 20 24 28 32 .4Angle of attack, a (degrees)

FIGURE 41.-N.A.C.A. 4418 airfoil.

chord

/.25 376 -2//

7.5 6.06 -4.67

20 //.72 -55625 12.40 -5.4930 12.76 -5.2640 12.70 -4.70s0 /1.65 -4023.60 10.44 -2470 6.55 -2,4580 6.22 -/.6790 346 - .9395 /.89 - .55

Slope ofrod)u5D through Z'endof

5/a. Up •r. Lm,o l0° - ° us 0

2.5 Soo -299 W k _/05.0 6.75 -4.06 C o

15 10.66 -5.49 Per centofchord/0 9.// -506 0 20 40 60 80 /0

4.E.Rad.:3.56 C If

c. .

0 L/D

n

Airfoil. N.A.CA. 4418 R.N.:3,100.000Size: 5"x30" ve/(f//sec.): 69.5

I-rl' Dcle:8-27-3/Where tested: L.MA.L. Test: l!0.T 655Corrected for tunnel-wall effect

aO

k2W

28 024 0 20

q 2004N ^

/6^ 6

o /2r8u

8 0/0`• C0 4 v4 u

0 4c

U

-8

0

0

0

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

25

FIGURE 40.-N. A.C.A. 4415 airfoil.

48

44

40h

36 w

32 v

28 Z{g

24 0l

20 m

168a v

2.'ck

4^

Oa

C ` -4 o

-B

-/2I: NA.CA. 4418 R.N.: 3,/00,0009-27-31 Test: VD. T. 655 -16cted to infinite aspect ratio.4 .6 .8 LO L2 1.4 /.6 /.8

Lift coefficient, C

270770-35-4

26 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

zo

D

t`o

v28 0

3241,20

Q 20x40y ^o /6^ 60

b

0/280

8 0/00

° 4 u00 Q

C-4 0.

-B

1252.550]5/B15253050BO706090

100l0 0

2013.174014.16

4.455.B47.82924%3512.0413 B614.2713. lB//.609.506.9/3B5

/.-

-2.42-3.46-4.78-5.62-6.15-6.75-6.9B.-6.92-s.76-6./6-5.34-4.40-•235-23/-/.27

F00

L^ r D 10g 0

W w _/0h p

0 20 40 60 80 lPer cent ofchord

LG Bod.: 4.65Slope ofradiusthrough end ofchord: 4/20

C

.c. pr-

L/D

C

Airfoil NA.CA. 442/ R.11:3.110,000Size: 5"x30" Ve/(H/sec.): 69.4Pres.(stird ofm): 20B Dote:8-28-31Where tested: L.M. A.L. Test: VD.T 656Corrected for tunnel-wall effect

N

B v . /s 0 .02

z

./2-

48

44

O

0

00 ./0 r 40n y

.44 .09-- -36 v8 ^

2.0 .40V .08 a; 32 U

1.8 .36 u.07 26 e

1.6 .32 0.06 24 0V U ^

/4 .28'^ 005 20 uv ^ Q

/.2V.24 y v.04 168

/.'O w .20 um °.03 12U R ^

° o 0.6 u ./2 .0/ 4

S uk.4".08 0 0

0 0 -.2 -8 u'u e

-.4•4 Dote. B-28-31 Test: VD.T. 656 -16.2

w '3 Air foil, N. A. A. 4421 R.N:3,110,000 /2

R- Corrected to infinite aspectrotio-8 -4 0 4 8 12 16 20 24 28 32 -4 72 0 .2 .4 .6 .8 LO /.2 1.4

Angle of attack, a (degrees) Lift coefficlent,CFIGURE 42.-N.A.C.A. 4421 airfoil.

0U

W

28 p 0

24 0 20k

Q 20F40V ^g` /6't 60

a 12- BO

8,0/00k c20 4.o

u

0 Qc

-4 0._8 d

25

7.5 3.25 - .97

20545 -:3225 59B .0230 6.36 .3440 6.74 ,9350 6.65 /.3560 613 /.5670 5.21 /.5360 390 /.25

290 /8 .7295 1.17 35

i°o°o /- /-0sl

5f. UPr.G'

'

r. D 10/. -. UC

2.5

.667/U _/0

05.0 260 -/.00 h p

/0 362 -.89 0 20 40 60 80/5 4.74 -.64 Per centofchord

L.E. Rod.: 0.4051ope frood usthrough entl ofchord: 4125

L D Cc. p.

0

Airfoil: NA.C.A. 4506 R.N,:$05Q000Size: S"x30" Ve%(H./sec.)Pres.(st'n:

70.0d o

M.tm.): 20.6 Dafe: 4-/5-31

Where fesfed• L.A.L. Test: V.O.T.568Corrected for tunne/-/l effect

e 44

0040

36-0.44 .090 0^2.0 .40 -08 ro 32 O

1.8 .36 .07 28/.6 .32^ x.06 24:944 .28, 0.05 '01

y N

/.2V.24^ y.04 /60k

/.0.1.200u 0.03 uo 12'u

.81./60 .02

.6 u .12^ .0/,o

w 4u.4 .08 0 0

r °

vO O y -.2 -8 ^`

°u e

•2 c •3 Airfo#-N..A.C.A. 4506 R.N:3.050,000 -l2-.4 - 4 Oote: 4-/5-3/ Test: V AT 568 _l6Corrected 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 /.8Angle of otfock, o: (degrees) Lifl coefficient.FIGURE 43. N.A.C.A. 4803airfoil.

51./.252.55.07.5/0/5202530,40506070809095100

/00L. E.Slopethroughchord:4

up',13S2.473.544.36SOS6.126.897.467.856./97.977.276./24.56256/.(.09)-Rod.:afrad'us

end

V0,-1.12-1.50-1.84-1.99-2.05-/.96-1.75-/.47-1.16- .52

.03

.42

.62.60.35/5f-.0910

0.69of

25

D /OU ^ 0L, -/0

0 20 40 60 80 /0Per cent ofcho^d

C

LID

i

C

Airfoil, N.A.C.A. 4509 R.N..:3,120.000Size: 5"x30" V.I.(fG/sec.): 69./P es.(s fnd. almf:20.9 Dote: 4-15-31Where fested: L.MA.L. Test: V D.T. 569Corrected for tunnel-walleffect

a0Uk

v

28O324 0 2iw

N 20041

p /6 61D o°12^Bi

8 0 /Oic

° Suov^ 0 Q.0

-4 0.

G-B

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 27

.44

2.0 .40

/.e .36

/.6 .32

V/.4 .28c

,W

/.2V .24v

LOv.20^

8x./60o O

.6h ./2

.4'.08

.2 .04

O 0

-.2

-.4-8 -4 0 4 8 12 16 20 24 28 32

Angle of attack, a (degrees) Lift coefffcient,Q

Ul°tU

aCv

28 0

24 3L. 2^o

X20041N ^o /6^ 61

o /2 Cu

e

B o /Oi

0 v4o u

v^ O4C

4 0.v-8

-0 -4

Proves 44.-N.A.C.A. 4608 airfoil.

.44

2.0 .40

48 .361.6 .32

Ci/.4 .281

W2,j. 24 i?

/.0 r .20 0u

.6 .1600 0

.6°./2k

4".08

.2 .04

0 0

-.2

-.4

O 4 8 12 /6 20 24 28 32Angle of attack a (degrees)

F'lovac 46.-N.A.C.A. 4512 airfoil.

5o.12525SO7.5/0152025304050

709095/00

/00

606.43805.23

Up',2.333.224.505.466.257.46a348.959.379.649.29

7.062.93/.58(./2)-

L'-r.

-2.07-2.67-2.99-3.18-3.28-3./7-2.95-266-/-/.97

.29- .70-.2,9-:04

.00- .04(-. 12)0

vD /0OO0 20 40 60 60 10

Per cont ofchord

CLX. Rod: 1.56Slope ofrodusthrough a 2Zchord: 4/25

LID C.P.

r

AlrfOlL'NA.CA.45/2 R.N:320Q000.Size: 5'k30" Ve1.(fL/sec): 68.2Pres.(sfnd ofm):2/./ Dafe:/2-/6-3/Where tesfed'LAIA.L. Test:V.DT 733Corrected for tunnel-wall effect

Lift coefficient,C

48

44

40

N36 w

a32 O

28 ZS

.d24 p

20 u

/6gv

12 1J6 ^^

k4y

40

-4 0v

-B

28 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

D

s°

OcN

2B o 0

24 0 20ti

y 20040l

/6^ 60

o /2c80

8,0100

4 C0 v.o u

04C

-4 R

v-8

20/0ow o .12u ^

7

St..0

/.252.45

07.5toIS2025

6070BO90

100

3010.804011.019501062

Up".-

2.94

4 495.6.60].4s8.859.6//0.47

9.56299SB9332

-

L'w^-.o

-/.88B

-300 -2464

8-398-4.30-4.59-450

-4.16-3.44-262-1.e5

0. o

0 20 40 60 BO 10Per cen/ofchord

-1.20- .69- .35

O Ci

L. E. Rod.: 246Slope o/ radiusI= 9,5 end ofchord: 4125

C.P.

LID

C

Airfoi/-NA.CA.45/5 R.N.:3,080,000Size: 5"x30" Vel tft./sec.): 69.6P es.(sfndofm.):20.7 Dote:4-/6-3/Where tested'L.MA.L. Test: VD.T. 571Corrected for funnel-wall effect

.44 .09ti

2.0 .40 •x.08v

48 .36 x.07L6 .32 x.06

v/.4 .28- 0.05

va1.2V .24^ v.04

v

/.O.w.20' x.03U

1 h

.B m'/6^ .02o '0

.6 u .12 .0/w

.4JY

.08 O

.2 .04 V-./O 0 w -.2

U

-.4 0-.4-8 -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,CFcaysE 46.-N.A.C.A. 4518 airfoil.

skav

26 0 024,b 20

O 20 0 40y ^/6^ 60

o /280

60/o0`^ C0 4vo u

04c

-4 0.

-8

'r.

u

Slo.

0/.252550/0203040506070609095

/00/00

/5/0.25

2511.96

UP-

3.564.796.49

6.74//.27

12.37/2531/.94/0.726.926573 692.01(.19)-

LWr

0-2.25-316-4427

-5.4/-569-6.01-5. '9'-5.66-4.69-3.94-296311-1.34- .7/- .44(-./9)0

y a cua

/0

l 04 0 0

0 20 40 60 80 ICPer centafchord

L. E. R. d.: 3.56d.Slope of-d,,,

c hord 4125of

c.p.

IL

C

•Airfoil N.A.CA. 45/6 R.N:3130.000Size: 5"x30" 3e).(ff./sec.): 66.9Pres. (sthd afm): 2L0 Dote:4Wherefested.-L.MA.L. Test.VOT 572Corrected for tunnel-wc ,H effect

.//

I

44 .09 - -

2.0 .40 .x.08--_.._ ..._.......v Cno

/B .36 y.07k

1.6 .32 0.06C

1.4 .28`c 0.05W a

L2^.24X v.04v

1.0 .a .20 4.03U

.8v./6o 02

.6 0

.120 .0/ a'.4 .08 0

.2 .04 V`-./

0 0 m -.20.2 .3 Airfoi/: NA.CA. 4518 R.N: 3 /30.000

_ _ Dote: 4-16-31 Test, VD. T. 572.4 .4 Corrected fo inflhVie ospectrofio

40

36 wl

32 a28 ti

168

k6i

4yu

40

40m

e

-8 -4 0 4 8 12 16 20 24 28 32 -4 -2 0 .2 .4 .6 .8 /.0 L2 14 /.6 LAngle of otfoch, a (degrees) Lift coefficient, 0

Frooxx 47 .-N.A.C.A. 4518 airfoil.

S

0

m

26 O024 t. 20

kQ20E

o 40y

16 -60L 00/2280

60100c

0 4u

0 Qc

-4

-6

sl.

1.252SSO7.5/020

40506070809095

/00100

1511.632513.4630'3.06

UP'r.

4.235.607500.90/0.00

1273

139713.2711.869.657.264062.24

(22)-

L'v, r.

-2.56-3.B6-505-5.90-6.50-7.17-7.42- 7.38- Z/6-635-5.27-4. /2-3.00-1.97- /.05- .63(-22)0

U /O

u i 0

94MmN k -/Op

0 20 40 60 80 /CPer cent .{chord

LE.Rod.: 4.855/opeofrodiusthrough end ofchord: 4/25

Ci

c.p.

L/D

Co

Airfoi/: NA.C.A. 452/ R.N.:3,150,0006/ie: S° x30" Ve/.(ft./sec.J: 69.0Pres.(sfnd.otm).• 20.6 Dote: 4-/7-3/Where tested:L.MA.L. Test: V.D.T.573Corrected for tunnel-wall effect _4

0

-6 -4 0 4 8 /2 16 20 24 28 32Angle of otlock, a (degrees)

FIGURE 48.-N.A.C.A. 45

7

.44

2.0 .40W

L8 .36 o

1.6 .32 0

Ci u

1.4 .2Bc 0v $

L2 V .24 v

v 0LO 0u

ev.160o O

.6^./2

.4 .08

2 .04

0 0 y

ou.2

c

//j -4

,44

2.0 .40CN

/.8 .36 .0

/.6 .32 0V u

1.4 .28 C 0N

L2V .24w vv '^

/.0_v .20 uu p,

.Bo .lso

.6' ./2w

.4'.08

J.2 .04 VE

0 0 y0

-.2_4 0

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE -DENSITY WIND TUNNEL 29

/2 48

.//

E44

Vco./0P 40

.09 0 311

.08 hl

32 Do,.07 28 tt06

024 p

l.05- zo

v

.04 p16

.03 2

02 ° B0/

0

qtu0

00

0C p -40

_2 v-80

e.3

-.4Airfoi..• N.A.CA. 4521 R.N3: /50,000Dote: 4-17-31 Test: VD.T 573Corrected to infinite ospect rotib

/2_16

.4 .2 0 .2 .4 .6 .8 LO 1.2 /.4 t.6 L8Lift coeff/cient,0

toll.

OLkav

28 0 03

24 0 20

4 20 a 40

g16D60

o /280

8 0 /00

4 uo u

0 Qc

-4 0.

- 8 u

S^./.252.5SO7,5/015202530405060706090100/o01-T ROd:/7_6T_

Up ^.2.263./34.36S.6.027.176.0/8.6090l9369.166.567.445.66323!./2)-

Gwr.-1.57-2./6-2.81-3.17-3.40-3.56-3.50-331-3.00-2.27-/.4/- .56

./0.40.3/(-.%210

QD /0 ITTi 0

-/OO 20 40

Per cent60 60 /C

ofchord

0

5/ape cf "ofthrough end ofchord: 4/30

c.p.L/D

a

Airfoi/.• N.A.C.A.46/2 R.N.:3,2/0,000Si­ 5"x30" Vel.(fl./sec):68.3P es.(slnd.olro):21,0 Dole: /2-/7-3/Where tested.-LM A.L. Test: V.D.T. 734Corrected for tunnel-wall effect

X140

+-H 24

16

8

4

0

Airfoi/:NA.C.A. 4612 R:N:3,210,000Dote: 12-17-31 Test: V D. T. 734 -l(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 CcF-it. 49.-N . A.C.A 4612 airfoil.

44

2.0 .40 UN

18 .36 .0

16 .32 0V U

/4 .28cv ^

A2,j. 24 bv '^

/.Oy.20°u ^4

0 0.6 .12

.4 .08

.2 .04 y`

0 0 yu

2c

E_.4

-8 -4 0 4 B /2 /6 20 24 20 32 £

l0 5.64 -3.56

20774 -377256.3/ -3.6030 9.71 -15640 9.06 -2.5650 6. - /.6460 6.4747 - .6570].6] .34BO 6.21 .9590 3.7395 .4202)

100 L/2) (-/00 0

0

St . up L--

No /0/.25 221 -/.6/ U t 0253.04 -222 k -/050 4.24 -292 y °10 5.13 -3.32 O 20 40 60 80 /L/56.95-3. 79 Per ceniofchord

Ci

chord.

.78

L. E. Rod.: 1.56Slope ofrodLsihraugh4/3sendof

n

Airfoil: NA.C.A. 4712 R.N:3,160.000Si-S"x30" 1/e/(tf./sec): 68.7Pres. (sind.oim):210 Dote:12-18-31Where fested'L.MA.L. Test: V.D.T.735Correctedfor iunnel-wo//effect

0U

v

28 D 0

324 0 20

20 0 40l

p 16 ^ 60L

o / 0t 80k

6 °loo0 40.0 U

v

C4

0.

-8

7.5 6.42 - .46

20 //.74 .2625 //.92 .033011

- . /O4011. -./750 /0.49 :.1260 9.// .0570 7.37 .0/80 5.30 .0290 2.99 - ,03

loo (,/2) (-.",/00 - O

Sfa. up'r. L'0, O cu0 - 0 y /O

/.2S 370- .79 U l 02.5 4.99 - .94 y M. 1050 6.92 - ,79 y p

/s //.09 Per centofchoro/0 9.56- ,15 0 20 40 60 80 10

95 1.55 -:03 ( L

L E. Rad.: 15B5lope ofrodi3sihro39h end ofchard: 6110

c.LID.

n

Airfoil: N.A.C.A. 6212 R.N.:,^240,000Size: S"x30" Vel.(ft./sec.): 66. lPres.(st'nd. ofm):20.9 Dote:12-21-31Where tested: L.M.A. L. Test: V.O.T. 737Corrected for funnel-wall effect

°C

Oav

28 p 0

24 kk 20^o

V20 40

o /6. 60

12,80

8 0 /00

° 4u

0 4c

-4 0.

-8

30 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

4844

0 40

09 36

o.08 32

.07 28

.06 24

°.OS 20

.04-/6

.03 ao /2

.02 8

.0/ 4

0 0

-4

-8

.3 Z- toil. IV. A.0.A. 4712 R.N.: 3, 160, 000 /2_ Dote: 12-18-31 rest: VD. T. 735

' 4 Corrected to infhfte aspect ratio lB-2 0 .2 .4 .R A CO 12 14 1R 4R

Angle of ottoc/, or (degrees) Lift coe fficienf, CFIGURE 50.-N.A.C.A. 4712 airfoil.

7

.44

2.0 .40 GC

/.8 .36

1.6 .32 0V

1.4 .28 0N ^

L2V.24- v

v40i .20 ou h

.8 tE ./6o

.6u

.4". 06

.2 .04 Ci

0 0 yu

_40

-8 -4 0 4 8 12 /6 20 24 28 32 c`Angle of offock, a (degrees)

./2 48

44

0 40

Co, ;.09--36

0.08 32,

.07 28

.06 24'

°.O5 20

04 /6

.03 /2

.02a

8

.0/ 4

0 0

-41

_.2 _6

.3 Airfoil: N.A.CA. 6212 R.N. 3,240.000 -/2-4- Dote: /2-21-3/Test: V D.T 737 _16Corrected to infinite ospecf ratio-4 -2 0 .2 .4 .6 .8 LO l.2 /.4 1.6 /.8

Lift coefficient, CFIGURE 51 .-N.A.C.A. 6212 airfoil.

sf..1252.55075to/520253040506070809095

100l00

UP r1.632463.794.67SB07.268.238.809.006.766.177205.904.272.341.24(06)-

L r.- .40-.34

.OS.50

.96/.792.452.643.002.982.862.622.211,63.86.43

(-.06)O

v ^ /Ou 0ly 0 -/O

0 20 40 60 80 100Per centofchoro

0L.E.Rod.: 0.405/ape of'oo-lhra3qh end ofchord: 6//5

LID C.P.

C

Airfoil: N.A.C.A. 6306 R.N:,$080,000Size: S"x30" Ile%(fr/secf:69.8P es. (st'nd. otm.):20.6 Dofe: 4-/7-3/Where tested.• L.M.A.L. Tesf: VD.T. 575Corrected for tunne%wol/effect

uk0

v28 0 024 k 20

q 20040

0 /6 60

/2U 80

B 0/000 4 W

.o u

v02

C-4 0.

-8

./O

.44 .098

2.0 .40 x'-.08

48 .36 ^.071.6 .32 0'.06--

G U1.4 .2B c 0.05.v/2V.24-Zu v.04

v `^40y.20, k'.03

u O,.8v.16o .02o a

.6.12 .0/w

.4 J .08 0g

.2 .04

0 O -.2

-.2 -.3c-.4-.4

-8 -4 0 4 8 /2 16 20 24 28 32 -4 -,Ang/e of aHo'k. a (degrees)

FIGURE 62,-N.A.C.A. 6306 airfoil.

Slo.O

/.250

-.75

20

/O

U2

255.0r0S 767.5/0/5

50./B.46

0204060BO/OPer Gent ofchord 0 ./0

.44 .098

2.0 .40fW

L8 .36 •u.07

/.6 .32 '0.060 u1.4 .28-, p.05

v$/. 2V.24 ,4y v.04

v '^/Ov.20a x.03U ^^

.02O ^

.6i .12 .0/

.4^ .08 0

.2 .04 V -./

0 0 y -.2OV-.2 ^-.3

-.4 0 -.4-8 -4 0 4 B 12 16 20 24 28 32 `C -.4

Angle of otlock a (degrees)FIUVSE 63.-N.A.C.A. 6308 airfoil.

SUaN

28 0 0324 0 20

q20 0 40

p 16 D 60l lo /2 ^ 80o

80)00

0 4 u0v^ OQC

40.U

-8

2025304050607060909095/00

looL.E.Shope/hroughchord.'

9.7010.2910.5010.249.506.356334.942.71/.45

cos)-

Rod.:ofrad'us

end

1.021.351.501.531.55 /.4B/29..5050

23!-0

o0.89

of15

C

C.P.L/D

C

Airfoil: NA.CA.6309 R.h!:3.110,000Size: 5"x30" Vel.(R./sec.): 69.4Pres.(sihd.olmJ:20.B Dole:4-18-31Where fested: L.MA.L. Test: V.D.T.576Correc led for funnel-wo//effect

y36

32

28 tt

20 w

/6 0

12'k

k4^

00p

-4 0ro

-8

40

h360l

32

28 tl

160vc

6CO

4^u

0:,L,0

-4 0v

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 31

AirfOi/:NA.CA. 6306 R.N.:3,080,000Dote:4-17-31 Test: V D.T. 575Corrected to infinite aspect ratio

O .2 .4 .6 .8 /.0 /,2 44 /.6 /.Lift coefficient,C

4-18-31 Test: V. T. 576Gted to fnfnile aspect ratio.4 ".6 .8 /.0 1.2 /.4 16 1

Lift coeffi'cient,C

40

Ll 7.26 - /.47

20 //. 14 -4025 /1.79 - . 1430 /2.00 .00

11.69 .08

50 10.84 .2360 9.50 .3570 7.76 .3990 5.62 .3490 3.08 ./S95

1.66 .03

St... Up. L'w'r. 1b /0

/.25 3.05 -6---

U 025 4.20 -1.38 yk y05.0 5.93 -1.56 4 0/0 9.36 -/. 29 0 20 40 60 80 /0015 /0.03 - .83 Per cenl of chord

/00l00 (./2) (-./?j C

G. C. Rd.: 158S/ape ofrod'us

yn endofchord: 6/I5

L/DC.p.

Ile

C

Airfoil: N.A.C.A. 6312 R.N.:3.f80,000size: 5"x30" Ve1.(R../sec.): 68.5Pres.(sfnd.otm):2/.0 Dote:12-22-31Where lested: L.MA.L. Test: V.D.T. 738Corrected for tunnel-wa//effect

DO

o0

aW

28 0 03

24^ 20

q200 40

0o /sa 60

o /2 u 88'-/00

c0 4vO U

v04-4 0.

-8

'r

u

0

.44U§2.0 .40c

/.8 .36/.6 .32 0

V u1.4 .28 o

v1.2x.24

v k1.0 ,v .20

U6^./6 0O O

.6X./2k

.4 .08

.2 .04 G

0 0 m

U2

cF_4 c

32 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

44G$

2.0 .40W

1.8 .36 w

1.6 .32 0U

1.4 .2,9-, 0

/.2V.24' wv '^

40 .208 o

e`* ./s o0 0

.6X ./2w

.4J.08

.2 .04 G

0 0 wou

.2cv

-.4 0-6 -4 O 4 8 12 /6 20 24 28 32 F

Angle of attack, o: (degrees)FIOVAB 64.-N.A.C.A. e:

./2 48o x est No. 738.// oo q 577 44

do 40

09 36

.08 32,

.07 28

06 24'

.05 20'

.04 /6

.03 /2:

.02 8'

.O/ 4'

0 0-4'

_2 C. -8

' 3Airfoil: N.A.CA. 6312 R.N.:3,180,000 -/2_ Dole: /2-22 3/ Test: V. 0. 738

' 4Corrected to infinile ospect rolio -/B-4 -2 n .2 .4 .6 .,9 LO L2 14 L6 1.8

Lift coefficient.0W.H.

u

ll

i

U

0

'w

v28 0

3

24 0 20k

020;4^p /6 6

l l

o l2r 8u8 0 /0

0 4 vo u

0 4C

-4 4

_8 v

D

0

-8 -4 0 4 8 12 16 20 24 28 32 -4Angle of allocic, w (degrees)

FIGURE 56.-N.A.C.A. 8316.11U1.

Sic. up'r. LM,0

U/0

/.844

t00

/0 9.6] -236 0 20 40 60 BO 10/5 //.45 -2.13 Per ceni ,fcha d

/25 399

"375 948 -2.39

20 /2.60 -1.6325 /3.25 - 1.6230 /3.50 -/.5040 13.15 -/.375012.16-60 r0.67 - .77]O 8.7/ -.50BO 6.30 - .3090 a45 - .2095 /.BJ-./J/00 (./6J (-

L.E. Rod.: 2485/ape ofrod'us

0 through endofchord.' 6//5

25 5.15 -4SO ].OS -2.27

_/0

0

/00 - 0 C

c. p.

o L/D

0

C

Airfoil. N.A. C.A. 6315 R.N:3,100,000Size: 5"x30" 3el(0.1sec.): 69.7

_ Pres. (surd almJ:20.5 Dote:4-20-31Where tesled.- L.M.A.L. Test: V DT 578_ Corrected for tunnel-wo/l effect

40

36

o° 32

28

24

20

/6

6

a° 4

0

4-8

/2Airfoil. NA.CA. 6315 R. N: 3100,000Dole: 4-20-3/ Test: V D.T. 578 -16Corrected to inf nice ospect ratio2 4 R R /0 /2 14 /R IR

Lift coefficient. C

'z.e2 .44 .09

2.0 .40 ..08e

vLB .36

16 .32 0.06U

1.4 .28 0,05my

L2V .24'M x.04v `^

l.Ov,208 11.03--U 4

.8^.16^ .020 0

6X .12 .0/k

.4' .08 0Y

2 .04k _0 O U 2

_ ' 2 .3

--4.4 .4-.4

56.-N.A.C.A. 6318 airfoll.

1.40455.36

-./9l.5aY'us

fC

5

c.R

L/D

n

Airfoi/: NAC.A.6318 R.N:3080,000Size:5"z30" V.Z(ft/sec.): 69.3P es. (sfnd.otm):2/.0 Dofe:4-20-3/Where tested: LMA.L. Test: VD.7.579Corrected for funnel-//effect48 12 16 20 24 28 32

_?91' of attack. a (degrees'Fcccaa

h

vU

Da0

0

160v

l2ck

4

4y

U

030

-4 0v

_/2e

Dole: 4-20-31 rest: V. D. T. 579Corrected to k7finite aspect ratio l

.2 .4 .6 .8 l.0 1.2 1.4 1.6 1.8Lift coe ffcienf,C

O0V

0D

v

28 0

324 a

kq 20 0i0 16l

0 /2

90

0 4 N.0 U

WD CC

-40.Bu

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

w o /0 2u ^ O

o -/0

0 20 4060 60 /00Per cent ofchord 1 ./0

D

tiaC

28 0 I

24 wa 2

q 20x4y ^o 16D6C ^o /2 B

8010

0 4:o uCv0 Q

C-4 0.

-8

Ito./.252.55.07.5l0l520253040506070809095

%00

Up'r.5.707.209.3611.0312.3514.2615.54/62416.5016.0614.6412.99/0607.694.222.31( 22'

L'Owr.-1.79-2.65-363-4./5-444-454-4.65-4.58-450-4.25-3.7/-3.02-230-1.60- .90-.55( 02)

1, o /0u S Ol U0 -/0

0 20 40 60 80 /0Per cent ofchord

L. E. Rod.: 4.855/ape ofrad LSthrough end ofchord: 6//5

C

cp

i

LID

C

Airfoil.•IV.A.CA.6321 R./V.:3/40.000Si- 5'x30'Uel(ff./sec.): 69./Pres. (stird.otm):20B Dof-4-21-31Wherefesd:LMA.L. rest: V. D.r5B0Corrected

fe for tunnel-wol/effect

.//

.44 .05G$

2.0 .40 •.06

1.8 .36 . v0.07

1.6 .32 0.06's

1.4 .28-, 0.05va

420.240 m.04m `^

10y.20 .03.0

.BV./6o .020 0

.6,U .12 .0/

.4' .08 0

.2 .04

0 0 0 -.2u-.2 ^ -.3v

-.4 0-.4

h36%

l

32 D

28 21d

24l

20-.v

/6 pv

2c

8c0

4^

0o^0

-4 0w

-8",Q

AirfoiLNA.CA.6321 R.N:3,140,000Dote: 4-21-31 7ez f: V. D.T 580 1_

FTT-R Corrected to infinite aspect ratio4 -2 0 .2 .4 .6 .8 LO 1.2 /.4 l.6 1.8

Lift coefficient C-8 -4 0 4 6 12 16 20 24 2B 32 -.

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

7.5 4.24 -.06

20 7.42 1.6125 8.16 2./630 8.64 2.6240 6.90 3.1050 9.49 3.1960 7.64 3.05

Se. u,. -,. L'W^. zD /0Z25 /.45 -.52 02.5 2/6 -.55 L,4SO 3.32 -.36 y p-/0/0 506.2B O 20 40 60 80 /6/5 &J9.97 Per centofchord

70 6.35 2.66ao ass 2.ozsa 2.59 /. u95 /.3B 55/00 (.061 (-.061

100 - oL.E. Rod.: 0.40Slope If-d",Clhroughchord: 6/20120

C

Airfoil: N.A.C.A. 6406 R.N.:3,100,000Size: 5"x30 •' Vel. (ft./sec.J: 69.2Pres. (stud. ofm.): 2/.0 'Dote: 8-31-31Where tested.• L.hLA.L. Test V.O.T. 65BCorrected for tunnel-wolf effect

0

w

W

28 p 03

041, 20wq 20 o40

/6160

4^ 12 u 80

8 0 /00kc° 4 u.o °

0 4c

-4 0.-8 u

.2 .04 C0 O o

u-.2v

-.4-8 -4 0 4 8 12 16 20 24 28 32

44U$2.0 .40 c

/.8 .36

1.6 .32 0U

/.4 .28- 0•v a

1.2V .24v '^

LOV.20u^

A ./6 0

.6U./2

4^ .08

.44

2.0 .40

1.8 .36

1.6 .32d

/.4 .28 c.w

/.2V .24

40. .20 u

A./6./6O O

.6 ° .12

.4-'.08

.2 .04

0 O

-.2

-.4-8 -4 O 4 8 /2 16 20 24 28 32

Angle of offock, a (degrees)FILM. 69.-N.A.C.A. 64W .M.11.

5.0 4.30 -/./B

20 966 ./725 965 693010.13 %/240 /035 665SO 9.B/ 48660 8.78 1.9270 7.28 L7680 5.34 /.3690 2.95 .74

/00 /.091 (-.09)/00 O

S1a. up}. L'wY. C C /O

(25 206 -, 0 BB U•C O252.96-L// 40_/0

/0 5.4211 -/.66 O 20 40 60 80 /0/5 7.78 -.36 Per cenfofchord

95 /.57 35 C,

L.E. Rod.: 0.89S/ape ofrodiuslhrough end pfchord: sFzo

p.

Airfoil: NA.CA. 6409 R.N.:3,060,000Size: S"x30" 1/eL (f/./sec.): 698Pres.lst 3 otm):20.8 Date: 9-/-3/Where tested.• L.0 L. Test: l!D.T 659Corrected for tunnel-wo//effect

^pU

0

v28 p 0

324 0 20

420040

p /61

60

£'2 7 80

8 0 /00o 4m.o °

l

C 04c

-4u8

34

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

S

u

i

.12 48

44

./0 40

09 36

.0B 32.

.07 28

06 24'

.0$ 20

.04 /6

.03 /2'

.02 0 8

O/ 4'

0 0:

Airfoi/: N.A.CA. 6406 R.N: $ /00,000 -/2_ 4 Dote: 8-31-31 Test: U. D. T. 658 _16Corrected to infinite aspect ratio.4 -2 0 .2 .4 .6 .8 LO L2 /.4 1.6 1.8

Angle of offock, a (degrees) Lift coeffic/ent, 6iF-.. 66.-N.A.C.A. 6406 W.H.

9Yo./2525SO73to1S202530s06070BO909S

5/00

4011.110

up 'r.2733.605.366577.58.A18/03411.1411.65/1.169.956.23a333/.79/./2)

Uw'r.-123-1.64-/.99-,?.-1.99-1.67-1.25-.76- .38-.20

.55

.711

.65

.73

.39/6

('12)

C L /0U C 0h o_/0

0 20 40 60 80 100Per cent ofchord

CL.E. Rpd.: 1.511/ooe otrod/usthrough end of^hord:s/zo

C

Airfoih N.A.C.A.64/2 R.N:3,090,000Size: S VeG(f//sec.): 69.5

sf7dolm):20.8 Dote:12-22-31Where tested•L.MA.L. Test: V D.T. 739Corrected for tunnel-wall effect

.44

2.0 .40

LB .36 0

1.6 .32 0

Ci u1.4 .28c O

v^/.2.24 :0 v

ti

LO1.206U

e^./soo q

.6X./2k

.4 .08

2 .04 C

O 0 v

U_, 2eE_.4 0

O

u

Cv

28 0 0

24 20M

420E40y o

16 60

o /2g 80

8 `0/00k c0 4 m.O '

0 RC

-4 4

-B o

/6 pv

/2'c

8'11k

4s

0Oo

-4 ov

'B 0,e

-/2

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 35

Wv

Ua

-8 -4 0 4 8 /2 /6 20 24 28 32Angle of attack, a (degrees)

FIX 60. N.A.C.A. 6412airfoil.

rti

028 0 0

324 0 20

20 0 40y ^

16; 60l l

/2U 6

h 8 0/00` Co 4 vo u

0 Qc

-4 0.ti

-8

0

./2

0 0

44 .09

V$2.0 .40 - .08

N1.8 .36 .u.07

1.6 .32 0.06Co u1.4 .28 0.05

W

1.2V.24^E- k

x.04o

/.0 .0 .20 0.03U

.8'.16 0 .02o O

.6,', .12 .0/k

.4'.08 0

.2 .04 tj

0 0 Q, -. 2U

v-.4 0 -.4

-8 -4 0 4 8 12 16 20 24 28 32 -4Angle of attack, a (degrees)

Frovas 61.-N.A.C.A. 6416 airfoil.

z9 776 -ao6

20 /L6/ -26725 12.64 -2.2930 13.15 - L9140 1325 -1.2550 1246 - .76BO //. /O - .3470 9./6 -.0480 6.70 .0990 372 .0495 2.01 -.05

$YO. Up'r. L'wr.20

O - /. w p 10225 345 -/.53 -1 02.5 4.67 -2.13 4,4 0SO 6.44 -2. ]5 p

D

/0 8. so 0 20 40 60 BO /015 /0.556 -2.97 Per cent of chord

too 1'161 f-.16/C III

L. E. Rod.: 2.48Slope ofra0 usthrough end ofchord: 6/20

C. p.

L/D

a

Airfoil: NA.CA. 6415 R.N:3,060,000Si-5"x30" Vel(ft/sec.): 69.8Pres.(sthdotm):208 Dofe:9-3-3/Where fested.- L.MA.L. Test: V.D.T. 661Corrected for tunnel-wolf effect

2 48

00. Test No. 739•// 000 660./0 40

.09 36 v4

0.08 32 D

.07 28 t.0

.06 24 pl

0 .05 20,v04 16 p

.03 12 c

v

.02 8

.0/ ^ 4

0 0^

-4 0_, 2 C ^, _8 0

v

.3 Airfoi/NA.CA. 6412 R.N.: 3,0917,000 12.• -.4-- De:/2-22-31 Test: VD.T 739 _16

1co _ -cfed to infinite os ect rofb-4 .2 0 .2 .4 .6 .8 /.O l.2 L4 L6 L8

Lift coefficienl,C

Dote:9-3-3/ Test:Corrected to infinite aspect.2 .4 .6 .8 LO 12

Lift coefficient,C

/0

.44U^2.0 .40v

46 .36 .0

/.6 .32 0

U° U

1.4 .28-, o.v a

42V.24;^ vv k

44 .20 8.0 p a

Bv.16^0 a

.6X ./2

.4 .68a

2 .04 C

0 0 v02 c

e_4

h36 m

l32 U

28 b.0

24 pl

20 ua16,v

12:c`c

^° o4r

000

-4 0C

w_8 0

Q

7: NA.C.A. 6418 R.N: 3.100,000 29-4-31 Test. VO..T 662 _16tied to infinite aspect rcho

.4 .6 .8 10 12 1.4 L6 7.8Lift Coefficienf.0

36 y

32 D

28 8.d

24 0l

20 w

/6 0v

12:c8.c

44^

000

-4 00

_8 P

Q

V O. T. 663

36

REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

ukb

v

28 0 03

24 0 20

q20040o 16 O 60

o /2^8k

8 v l0

0 4 uv^ 0 a

- 4 0.8u

0

0

25

30

1. 4.26 -1.92

].5 9.99 -39/

20 /332 -4.0]25 14./7 -3.7514.64 -34040 14.70 -2.7050 13.60 -20560 1224 -44770 /0.1/ -.9460 7.40 - .54

90 4.12 -95 2.24 - .23100 (./91 (-./9)/00 0

0 0 ur /025 7.62 -2.59 v.; O3.0 7.53 -346 C pl0 10.06 -4./5 0 20 40 60 60 /005 1202 -4.26 Per cent oFchord

Slope ofrodius1hro3qh end ofchord: 6/20

L E. Rod.. 3.56 C

c.p.

L/D

n

Airfoil: NA.C.A.6418 R.N.:3.100,000Size: 5"x30" Ve%(R../sec.): 69.4Pres.(st'nd.ofm.): 20.6 Oate: 9-4-3/Where tesled: L M.A.L. Test: V.O.T 662Corrected for tunnel-wo/t effect

-8 -4 0 48 12 16 20 24 28 32 -.4

Angle of attach, a (degrees)FIGURE 62.-N.A.C.A. 6415.W.11.

U

SkU

W

28 p 024 0 2w

y20 0 4616-1-6

,l

a 12 ^ 8

B o t0C0 4.

.ov0 Q.0

4

U-8

0

0

0

tl

5.0 5.65 -4.167.5 10.24 -4.81

20 /4.79 -5.4925 1565 -52330 16.15 -4.9/40 /6.16 -41650 /5.14 -3.4060 13.44 -2.5970 11.06 -/.9380 8.06 -1.1790 4.51 - .65'95 246 - .42loo (.221 (-.221

/00 O

Sta. up". L'w 'r. `0 - O y o /0

/.25 5.13 -208 Ur 025 6.60 -3.04 m 0 _/O

/O 1/.52 -5.18 0 20 40 60 80 t0/5 /344 -552 Per cent ofchord

L.E. Rad.: 4.85

Thro end ofchord: 6/20

Slope ofughradiuI C

c. p.

0 L/D

C

Airfoil: NA.C.A. 642/ R.N.:3,030,00014EE: 5"x30" Ve1.(fl./sec.): 70.5Prss.(sf'7d.olm.):20.4 Ooie:9-4-31Where tested: L.M.A.L.. Test: V.OT. 663Corrected far lunnel-wall effect

.//

44 .09

2.0 .40 ­08-v

18 .36 .0 .07

AS .32 E3,06u

1.4 .28 1 0.05W

ka .04

V ^ vLOy.20 03

.8x./60 .0201 a.6w .12 .O/

.4'.08 0Y

.2 .04 G -. /kO 0 0 -.2u

.2 .3 Airfoit: N.A.CA. 6421-.4 a._ 4 Dole: .9-4 -31

F Corrected to infnits os

48

44

40

-8 -4 O 48 12 /6 20 24 28 32 -.4 -2 0 .2 .4 .6 .8 10

Angle of offpch, a (degrees) Lift coefficient. CFIGURE 63.-N.A.C.A. 6421 airfoil.

s .1.252.5507.5/01520253040506070809095100/00

Up ^.1.362.003013.654.585.766.747.498.066.668.658.066.695.172.90/.57(.06)-

G'0

Y'.- .60- .70-.62- .43- . l9.37.95/.512.032643.353.48321L521.43

73(-.06)0

D /OC 00 _l0

O 20 40 60 60 / 0Per centofoh-d

L.E. Rod.: 0.40,,, ,hendofchord: 6125 C

L D c. p.

C

Airfoi/:N.A.C.A.6506 R.K:,$17Q000Size: 5"x30" Ve/. (R./sec.): 68.5

d.Pres.(stnol-):2/./ Date:4-23-31Where tested: L.MA.L. Test. • V.O.T. 586Corrected for tunnel-wall effect

v

U

v

1,111 004k0 20

012004001660

12 80

60/00 4 u.v u

0 QC

-4 0.

-8 v

.44V2.0 .40W

l6 .36 u

16 .32 0

G °1.4 .28-, 0

N U

L2V.24 ^-y v`

/.Ov.20^ ou U^

6 /6°0 0

.6u./2

.4'.08x

.2 .04 C

0 0 y0-.2

v-.4 a

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 37

./2

0./O

44 .09

2.0 .40 Ci.09W

/.8 .36.07/.6 .32 0.06

V u44 28C 0.0542, .24^u \y',.04

v1.q.2004.03

.dU

.160 .02

.6U.12 .0/

.4'4 .08 0s

.2 .04 yr0 0 0 -.2u

-2 e-3

4-6 -4 0 4 8 /2 16 20 24 28 32 4

Angle of attack, a (degrees)FIOME 64.-N.A.C.A. 8608 airfoil.

36

32

?6 ^

?4 pl

?0

/6p

4,

lC ^ m8

-foil: N.A.C.A. 6506 R.N: $170,000 -/2te: 4-23-31 Test: V D.T. 586 _/8rrected to infinite aspect ratio.2 .4 .6 .8 /.0 1.2 /.4 L6 1.8

Lift coefficient,'C

OU0av

28 'b 0

324 0 20

V 20 a 40p /81t 60a /2 u 80

8"0 100

° 4uo UvOQ.c

-4,-8

s11.01.252.5

507.5

O//520304010.11506070809095

/00LT5/ape/hroughchord:

/ 00

up'r.-

/.932.753.59 -4454.975.627.188.223.569.979.207.815.853.29/.78(.09)Rod.:

6

ofrodiusendoi

L'w'r.0

- .99- 1.27-/.3445-L-.96- . 9

.5//.392.032.342.301.87/.0753(-.09)0

0.89

25

cv o /0t 0

C a 0O 20 40 60 60 /G

Per cenl ofchord

C

p.

C

Airfoil:NA.C.A.6509 R.N.:3,110.000Size: 5"x30" Ve1.0../sec.): 69.4Pres.(51'ld. 11-120.6 Dote: 422-3/Where tested.• I'll"L. Tes1:VD.T.585Corrected for funnel-wo/l eff ct

Dote:

/6

/2

8

4

0

-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.

46

44

40

136%

32

28 e.O24 pl

20 u0

/6 0v

/2.cv

k4yu

0x0-4 0

v-B Q

.44

2.0 .40

/.8 .36

/.6 .32CS

/.4 .281v

L2 Ci.24cvo-v

/.0 y.20°v

0 0.6 °./2k.4' .08

.2 .04

0 0

-.2

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

Angle of attack, a (degrees)FIOU88 87.-N.A.C.A. 8616 slRoll.

Lift coefficienf,C

2.5 4.40-2.33

20 11.14 -3,332S12.00 -29530 /259 -24940 13.00 -1.525012. 2- .6260/450 .0e70 9.69 .4980 7.22 .6090 405 .36

/00 - o

S^ OPr LW r.l0

125 3.26 -1.68 U 0S. 4,4 _/0

0 604 -3.04 p

15 297 -3.56 Per cent ofchordZ5 8.34-355 0 20 40 60 80 /L

/00 (:/5i (-JS) CiL.C. Rod.: 246t ough ^dofchord: 6/25

c. p.

LID,

n

Airfoi/..NA.CA.6515 R.N:3,100,000Size: S"x30" Ve1.(ff./sec.): 69.4Pres.(sthd.otm.J:20.11 Dofe: 4-22-3/Where lesled• L. .40 Test: V.DT. 583Correcfed for tunneFwof/effect

b0ZU0aCN

28 D 03

24 L.w

N20040

16 60^ to /2 80

8'.100

a° 4u.o

OQC

v-B

38 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

U

-°C

O

cW

28t D

24,420

^20i4o I6g 6

o BBolo

° 4u00 Q

C-4 0.

-8

D

00

0

tzs 2n -/..laRLI 1111

20

w -/o .// o a x Test Meoo

.44 .09d

2.0 .40 G.08C

1.8 .36 w.u.07

1.6 .32 0.06

V U1.4 .2B 0.05

•v D.2 24 x.04

vk110 4.03

a.8v.1.6 .020 0

.68 ./2 .0/w

.4'.08 0a

.2 .04w0 0 0

-.2-.2 u -.3

v-.4

-8 -4 0 4 8 /2 16 20 24 28 32 F -44ng/e of attack, a (degrees)

Fiaass 66.-N.A.C.A. 8613 ailfb 1.

5fa LPL-.a - o25 356 -/.82

/O 7.06 -2.45

20 9.69 -1-9125 /650 -1.473011.07- .984011. 6- .06so 229 7/60 1035 i.2170 876 L39BON6.5 I.2490.72

Ioo oL. E. ROd.: I.SB5roae ofrodisThrough end ofchord: 6/25

SO 5.02 -226 y o75 613 -2.43 O 20 40 60 80 /615 857 -2.27 Per cenf ofchord

L/D

c.p.

C,

Airfoil.• N.A.C.A.6512 R.N.:3,180.000size: 5'x30" Vel.(ft./sec.): 68.6P es.(sfndolm):20.9 Dote: /2 23-3/Where lesled: L.M.A.L. Test: V D.T.740Correcied for lunnel-11 effect

/2-23-31sled to infinite as.4 .6 .8 /.0

Lift coeff/cieni,C

w^ /ou °c 0C o -/0

0 20 40 60 60per cent ofchord

zss 44

2.0 .40

/.B .36

l.6 .32

14 .281w

1.2V .24m

/.0 4 .20u u,

0 4.6'./2.4'.08

.2 .04

0 0

_ 2

4

Ls4Los.4105

S0B1 C56dlls

fS

I Ic.P. I I I

Pf

Airfoi/: NA.C.A. 65/6 1?.1:3, IOQ000Size: 5"x30" Ve%(ft./sec.//: 693Pres.(sti]dolm):20.8 D&.%1_31Where tested.-L.M.A.L. Test:VD.T.Corrected for tunnel-wall effect

582

4 8 12 16 20 24 28 32

./2

EFFF48

44

0//,0,1

^

40

C 36

1 .08w

°p0

32

07 - ° 28

0.06 24

0.05 ZOax.04k

/6

^.03 /2

.02 0 8'

0/ 4

0 0:

v-.2o Cw-8

w '3_'4

Airfoi/.•N.A.CA. 65/6 R.N:3,/00,000Dote:4-21-31 Test., V.O. T. 582Corrected to infinite aspect ratio

-/2

-16-4 ..2 0 .2 .4 .6 .8 /0 /.2 /.4 /.6 t.8

a

U

vm

28 024 0

kN'0o

0 16^t ^

u

k 8 0/C0 4 v

o ul0 2C

-4 R-8 u

5^ ./.252SSO75/o

25

5015.70809095

%0000

/512.792014.103015.604015.63601181

UP'n4. ]56.176.209. ]010.93

/505

2714576.594.B52.67(22)

L'^ r.-222-3.26-450-524-57/-6./3-.6.14-59Z-553-444-3.27-220

^ a ^O•^1Q 0R 0-/0

O 20 40 60 80 /6Per cent ofchord

-/.30- .6B- .32- .23(:22)

L.E. Rod.: 4.65 C6/ope of radius1 / eo endofchord: 6/25

c.p.

L/D

Airfoil: N.A.CA. 6521 R.N.:3,080,000Size: 5"x30" Vel.(fi!/sec.): 69.8Pres.(sfnd.ofm):20.6 Dote: 4-21-31Where tested: LM.A.L. Test- VD.T 58/Corrected for tunnel-wall effect

OUk

z6

24 l 21

020041

/6^ &

o 12 1 Biu

B 0 /0,

0 4u.o

0 4r_4'0.-8 u

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 39

ng/e of a//ock, a (degrees) Oft coeff/clent,Cr'. 68.-N.A.C.A. 6618 airfoiL

/0

44 .09

2.0 .40 V^..06

/.8 .36 x.07/.6 32 0.06U° u/.4 .28 0.05

v/.2V .24 v.04

L0^.20U x.03u R

.02

.6 U .12 .01

4`1 .08 0

.2 .04

0 Q 0 -.2U_.2 C_3

4

-8 -4 0 4 8 12 16 20 24 28 32

-4 -

Angle of oiiock, a (degrees)Fmm. 68.-N.A.C.A. 6521 M doiL

N36 m

l32 a28

.o24 p

lED'.

/68-

04^u000

40v

DAirfoi/. N.A.C.A. 6521 R.IV:3,080,000ote: 4 2/ 3/Test: VAT 581

Corrected to infnite aspect ratio0 .2 .4 .6. 8 /.0 12 1.4 16 l

Lift coefficient.

Angle of oltock a (degrees)

F1Qv v 71.-N.A.C.A. 6712 e4(oll.

Litt coeff/'aent,C

.44

2.0 .40

/.8 .36

1.6 .32

1.9 .28 C,d

L2V.24v

/.O y.20

o rJ.6'./2

.4'.06

.2 .04

0 0

-.2

-.4-8 -4 0 4 8 l2 16 20 24. 28 32

Angle of ot/ock, a (degrees)Fl°II88 70.-N.A.C.A. 6812.11M.

°

Lift coefficient.

S"/.252.5So75/0152025

so/70BO9095

/00/00

3010.514011.1450//./3

UPr.2.453.414.785.646.70&//9.179.95

0.569.337.194./62.28(./L

L'11-1.43-L94-247-270-278-269-2.41-200-1.5/-.49

531.44L98/.90/./9.60

(-.l2/O

^ /0M 0C o

to

O 20 40 60 80 /0Per cent ofchord

C

L. C. ROd.: L58Slope Ifrpdius/h .- end ofchord: 6130

p.LID

ix IC

Airfoil: N.A.C.A. 66/2 R.N.:.1250,000Size: 5"x30" Vel.(N./sec.): 68.0Pres. Ws 7d.otm.):

M21.0 Dote:/2-28-3/

Where tested.• L.A.L. Test.• D.T. 741Corrected for funnel-wall effect

0U

cN

28 p 0

324 L.

Q

20 0 40

/6 60

o /2.g 80

q,-/00c0 4v

o uN

R 04C-4 0.Ci

-8

U0suti0UCv

28 p3

24 0k

N 20

0 16 ^O Oo /2 u

k 8°

46vOQC

-4 0.U

-8

44

40

36

.32V

'28-,v

V .24•wvv.20au p,v./60O ^u /2V

.08

.04

0

40 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

a0suk0Dcv

28 ^3

24 0k

q 20p

0/60a /2 u

6 0,

w

° 4 uo ^

0 4c

-4 0.

8v

L ifl coefficient, C,

6fo.

/.2526S.07.5/0/520253040506070809095tD0

too

w,/.421.652.302.542.692.872963.003.002962.792.04/.46.774/(.06)0

2.49-2.49

Lm,- q2-1.85-2.30-2.54-269-28]-2.96-3.00-3.00-2.96-2.79-204-/.46-.7]- .41(-06)0

u c° 0

ERRO 20 40 60 80 /0.

Per centofchord

L. E. ROd.: /.19

.

P

a

.A. 00068 RN.:3.090,000" M(H./sec.): 69.5tm.):20.8 Dote:4-8-3/

d.L.M.A.L. Test: l!D.T.556or funnel-wall effect

-a -4 V 4 c4 ca JcAngie of oHccA. a (degrees)

./2

./t

.t0

.44 .09

'.0 .40 1 .08W

!8 .36 ^.07

!6 .32 0.06^q U

1.4 .26 0.05,v

!2V.241^- m.04

LOv.20 0

.lsp .02.evU

o O.6 ./2 .0/

.4".08 0x

2 .04 G-./

0 0 0 -.2°u

' 2c .3

'•4-4.4

no

a

Cw

Air.Dote:Corrected

foil.4

N.A.C.A.-8 31

to0006B

infinite aspectR.N.:.X6

Test: Y 6ro:

-2 0 .2 .4

Lift.6

coefficient,.8 l.0 12

Q14

W

QDb

i00cN

28 ^ !3

24 a 2iv

y 2 i4t

o

80/0ik c

° 4u.o L

Oy.c

-4 R

-8 u

/6 8

/2

M4,^

0

-4 0v-6 a

e

556

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

41

cDu

-2.552.62

Per cent ofchord

.44

2.0 40

/.8 36

1.6 .32ti

/.4 .28`Cv

/.2 V .24

LO• w .20 u'u o•Bv.16o0 0

.6 0 ./2G4' 08

2 .04

0 0

-2_ 4

-296-3.00287-ass-2.1s--1.70/.22

_ '33-.061(oo. /o

c.p.

L/D C

0

Airfoil: N.A.C.A.0006T R.N.:3,180,000Size: 5"k30" uel.(ft./sec.): 68.7Pres.(sti7d.o/m):20.8 Doie:4-8-31Where fested•L.MA.L. Test:VD.r 554Corrected for tunnel-wall effect

0 4 8 12 /6 20 24 28 32Angle of attack. a (degrees)

Ftoo z72.-N.A.O.A. 0006T airfoil.

FlomtE 76.-N.A.O.A. 00068 abfolL

7

44 .09

2.0 .40 <.08W

1.8 .36

w16 .32 0.06

V U/.4 .28< 0.05

• a42U.24* -W.04

m '^1.01.20$ x.03

u u.8 0

.16 a

.02

.6'.12 .O/

.4.08 0--.x

.2 .04k0 0 0 -.2U

v-.4 1-.4---

-8 -4 O 4 8 12 16 20 24 28 32 -.4 .2Angle of attach, a (degrees)

Fioasa 74-N.A.C.A. 0012T SW0.

sfo.0125as5.07.5/O/520253040506070809095

/ooLao

upr.0

1.25L882663.6/4.215.095.64S.

6.005.755//

3392.43-,f.37

78(a2i0

4.29-4.29

cwr.0-12s

-188-gas-361-421-509-.464-5.92-600-5.75-5.//

-339.43

-137-.78

0

a "/o

l U4 0 -/0

0 20 40 60 80 /LPer cent olchord

L.E.ROd: 0.40

c.p.

L/D

C

Airfoih N.A.CA. 00127 R.N.:3.120,000Si- 5"x30" Ve/.(fL/sec.): 69.5P es.(sti+d.otm):20.5 Dole.' 4-3-31Where ies7ed L.M. A.L. Test: V D.T. 548Corrected for tunnel-walleffect

u

0

v28 0 0

24 ^ 20

V20 0 40

U 16 60

12$ BO

8 /00

0 4 V.o

0 y.c

_4 4ti

-8

f.6 .32 0U

1.4 .28-, oma

/.2V .24* vv ':

401 .20 0

.8x./600 0w

.4^ .08

.2 .04

0 0 moU

2 c_4

-8 -4 0 4 8 /2 16 20 24 26 32Angle of attach, a (degrees)

12525507510f52025304050.6070809095

100loo

ze43.694.595.08S.S.

5926.006.005935.594.984.072.911.55

63( ..12)0

-1 e4-3.69-4.59-5.06-536-574-5.92-600-6. 0-5.93-559--4.96

c l /o

0 -y

R a -10

0 20 40 60 80 1C

Per cent ofch.rd

4.07-2.91-1.55-83(-.. /2)

0L.E.Rod.: 360

L<

c.p.

Ca

Airfoil: N.A.C.A.00128 R.N.:3.060.000Size: 5"x30" Ve/.(ft..Isec.)_ 70./Pres.(sthd.otmf:20.3 Date: 4-3/Where tested: LM.A.L. Tesf: V. D.T.550Corrected for tunnel-waf/effect

0u0Dv

28 Do 024 0 20w20040

0 /6 D 60

a /2 u 80w8 0 /00

^. c0 4 u

v0 V.c

4 0.

-8 v

7

.44

V$2.0 .40 -W/.B .36 .0

42 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

48

44

401

36032 U

28 i0

24 l20 m/6p12^c

4^U

O^0

-4 0v

e

Dole: 4-3-3/ Test: U. D. T. 5481Corrected to inf'niie aspect ratio j-.2 .4 .6 .8 /.O 1.2 44 L6 1.8

Lift coefffcient.0

2 48

44

./0 40

.09CPo 36

.08 32

.07 28

t06 24

.05 20

.04 6

.03-- /2°.02 8'

.0/ 4

0C .ro 0.

-4

-.2 -8'

'3-4

A/rf0lL• NA.CA. 00/28 R.N.:3.080,000Dole: 4-4-3/ Tesf:V. D. T. 550Corrected to Infinite aspect rotio

-I2

_/6-.4 :2 0 .2 .4 .6 .8 40 /.2 44 l.8 1.8

Lifl coefficient,CRama 76.-N.A.C.A. 00128 airfoil.

Sro.

25SOZ5to/520253040506070B09095

l00/o0

Up'r

2834.285.4/6.321648.46B.889.006627676.44S093.65205/./7(./B)0

L'wr

-2.83-4.28-.£4/-6.32-7.64

-846-8.88-9.00-8.62-761-6.44-5.09- 365-205-1.17!-.7Bi0

UR p -/D

0 20 40 60Per centofchord

80 /O

L. E. Rad.: O.B9

Cc. p.

LID

C

Airfoi/NA.CA: 00/BT R. N:3,/50,000Size: S"x30" Ve%(ft./sec./: 69.2Pres.(sfndatin):20.6 Dofe:4-7-3/Where tested:L.MA.L. Test:VOT 552Corrected for tunnel-watt effect

U0tUo0cN

z8 ° 1

24021k

V 20o4i

o /6 &^ ta /2 t 8r U

8,/00 4 vO U

04C

4 0.u-8

^a

i^

0.0U4.oa

ZB O24 l 20

N 20 0 40

0 /6a 60

12. 80

006 0 /00- -a 4.o

p 4

' Cu

-6-8

30 sod40 8.8950 6.3660 Z4770 6.1180 4.3790 2.3295 1.24/0o (.78)

/00 0

-9.0-6.69-BAS-7.47-6./1-437-2.32-1.24(-.16)0

,44

2.0 .40 -

,N/6 ,36 U.

/,6 .32 0q U

V1.4 .28%

yL2^.24

40,^ .20P.e Q)

o O.6U./2

.4-.06

.2 .04 C0 0 y0

- 20

_ •4

L.E.Rad.:]. /5C

ap.

411L/D

C

Airfoi/: N.A.CA. 00188 R.N.:3,180,000Size: S'k30 1,Vel-(ft/sec): 692Prey. (sfnd. otm):20.3 Dote: 4 -6 -31Where tested.• L.M.A.L. Test: V.D.T 551Corrected for tunnel-wall effect

-4 0 4 6 12 16 20 24 28 32Angle of attack. a (degrees)

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

43

.44

2.0 .40

/.8 .36/.6 32

1.4 .281v

2,j. 04

v/.0120Uu a

8 0= ./6 v

0.6X./2

.4 .08

.2 .04

0 0

-.2

-.4-8 -4 0 4 8 12 /6 20 24 28 32

Angle of attack. a (degrees) Lift coefficient,CF/evas 78.-N.A.C.A. 081ST slrfolL

Frayse 77.-N.A.O.A. 0018B elrfolL

S''7.5 5.29 -3.17

20 Z56-325 791 -3.9730800 -40040 7.63 -a96506.73 -.86760 S49 -a66704. -3.2760- / -2.6490 1.26 -1.6395 .66 -.9510 (.0 (-.13)100 O -

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!h/ ou endofchord: O. 153

L.E.itod.: /.SBC

C.P.

L/D

0

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Ye'TIsec.): 69.4

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48

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-.4 0 -.4 Date: 2-26-32 Test: V. D. 765 _/6Corrected to inf/Hite aspect ratio

44 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

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CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 45

-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, CFIGURE 80.-N.A.C.A. 0012Fo and 0012F, aufoas.

PRECISION

A general discussion of the errors and correctionsinvolved in airfoil testing in the variable-density tunnelis included in reference 8. In connection with thisreport, it was hoped that a more specific discussion ofthe various sources of error and separate estimates of

the various errors might be given. However, after acareful study of all the measurements it becameapparent that practically all the errors may be regardedas accidental; that is, of the type the magnitude ofwhich may best be estimated from the dispersion of theresults of independent repeat measurements. Themajor portion of these errors is caused by insufficient

sensitivity of the balance and manometers, by thepersonal error involved in reading mean values of

slightly fluctuating quantities, and by the error due toslight surface imperfections in the model. The last isperhaps the most serious source of error. The modelswere carefully finished before each test, but the pres-ence of particles of hard foreign matter in the air streamtended to cause a slight pitting of the leading edge ofthe model during each test. This pitting was probablythe major source of error in connection with the earliertests, but it was reduced for the later tests when the

necessity of a more careful inspection of each modelwas appreciated. After a considerable period ofrunning the particles in the tunnel were found to be-come lodged, permitting this source of error to be

largely eliminated during the later tests. For this

report, however, the effect of the error from this sourcehas been minimized by repeating the tests of many ofthe airfoils, including all of the symmetrical seriesoriginally reported in reference 2.

The magnitude of all such accidental errors wasjudged from the results of repeat tests of manyairfoils, and from the results of approximately 25tests of one airfoil that were made periodically through-out the investigation to check the consistency of themeasurements. The accidental errors in the results

presented in this report are believed to be within the

limits indicated in the following table:

a t 0.15o

CLmax0.01

- 0.03Cm, f 0.003

CD1(CL=0) 0.0006V 1 -0.0002

CD0 (CL -1)0.0015

- 0.0008

In addition to the consideration of the accidentalerrors, all measurements were carefully analyzed toconsider possible sources of errors of the type that

would not be apparent from the dispersion of theresults of repeat tests. A rather large (approximately

1.5 percent) error of this type is present in all the air-

46 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

velocity measurements resulting from a reduction inthe apparent weight of the manometer liquid when thedensity of the air in the tunnel is raised to that cor-

responding to a pressure of 20 atmospheres. Theeffects of this error, however, are reduced by the pres-ence of another error in the air-velocity measurementsdue to the blocking effects of the model in the tunnel.The measured coefficients, obtained by dividing themeasured forces by %pV 2, as well as the derived coef-ficients are, of course, affected by errors in the air-velocity measurement. Aside from this source oferror, it is believed that only two other sources need

be considered: first, the deflection of the model andsupports under the air load; and second, the inter-ference of the airfoil supports on the airfoil. Theangle of attack and the moment coefficient are affectedby the deflection of the airfoil and supports. Theerror in angle of attack, which is proportional to

Cmcl,, was found to be approximately —0.1° for an

° Maximum thickness in per cent of chordF1aunE 81.—Variation of liftcurve slope with thickness.

airfoil having a moment coefficient of —0.075. Theerror from this source in the moment coefficient isinappreciable at zero lift, but at a lift coefficient of 1may amount to —0.001. The errors resulting fromthe support interference are more difficult to evaluate,but tests of airfoils with different support arrangementslead to the belief that they are within the limits indi-cated in the following table:

« ±0.05°

CLmax0.00

— 0.02Cmcl4

10.001

CDa(CL=0) f 0.00021 0.0000

CDa (CZ =1) ±0.0010

The tunnel-wall and induced-drag corrections ap-plied to obtain the airfoil section characteristics mightalso be treated as sources of systematic errors. Such

errors need not be considered, however, if the sectioncharacteristics are defined as the measured character-istics with certain calculated corrections applied.

Errors in the tunnel-wall corrections, however, shouldbe considered when the results from different windtunnels are compared. For consideration of theseerrors, the reader is referred to references 9 and 10.

For the purpose of comparing the results from differ-ent wind tunnels and of applying these results to air-planes in flight, it is also necessary to consider the

effects of air-stream turbulence. In air streams havingdifferent degrees of turbulence, the value of the

Reynolds Number cannot be considered as a sufficientmeasure of the effective dynamic scale of the flowThe airfoil characteristics presented in this report wereobtained at a value of the Reynolds Number of approx-imately 3,000,000, which corresponds roughly to the

Reynolds Number attained in flight by a medium-sized airplane flying near its stalling speed. Consid-

eration of the effects of the turbulence present in the

variable-density tunnel (see references 11 and 12) leads,however, to the belief that these results are morenearly directly applicable to the characteristics thatwould be obtained in flight at larger values of the

Reynolds Number.N./2

,aroma tSymmetrii al ai^ foiR! i I I 27 per rodlan

04 x 4%Vy + 6%

ro

0Camber position in fraction of chord

(Abscissa of maximum mean-line ordinate)FIGURE 82. —Variation of lifGcurve slope with camber. Results for 12 percent thick

airfoils.

DISCUSSION

The results of this investigation are here discussed

and analyzed to indicate the variation of the aero-dynamic characteristics with variations in thicknessand in mean-line form. For the analysis of the effectof thickness, test data from consecutive tests of airfoilshaving different thicknesses and the same mean-lineform are used. The analysis of the effect of the mean-line form is made with respect to consecutive tests of

airfoils of the same thickness (12 percent of the chord)and related mean-line forms. The results are com-pared, where possible, with the results predicted bythin-airfoil theory, a summary of which is presented

in the appendix.LIFT

Lift curve.—In the usual working range of an air-foil section the lift coefficient may be expressed as alinear function of the angle of attack

CL =ao (ao — aaa)

where an is the slope of the lift curve for the wing ofinfinite aspect ratio and a Le is the angle of attack at

zero lift.

lot

a.9t

u, XE v,

v oavv^g14.&

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 47

The variation of the lift-curve slope with thicknessis shown in figure 81. The points on the figure rep-resent the deduced slopes as measured in the angular

range of low profile drag. These results confirmprevious results (reference 1) in that they show the

lift-curve slope to decrease with increasing thickness.

The camber has very little effect on the slope, asindicated in figure 82, although a rearward movementof the position of the camber tends to decrease the

Comber position in froction of chordfAbscisso of maximum --line ordinate)

FIGURE 83.—Variation of angle of zero lift with camber. Points shown are for 12percent thick airfoils. Curves indicate general trends for the different thick-

nesses.

given mean line without altering the camber position.The theory also predicts an increased negative angle

as the position of the camber moves back along thechord. The experimental values are compared withthe theoretical values in figures 83 and 84. The ex-

704 8 12 16 20 24Maximum thickness in percent of chord

Francs 84.—Variation of angle of zero lift with thickness. Numbers refer to mean-camber designation.

0030240 25

0 430 440 45

o630 640 65

63a

`64

ya,

v

i65

/.f

/. E

/.5

/.e

/.t

CL_.t

I

b

° 'Symmefrfcol series

2% mean camber0 23 seriesx 24 ••+ 25

4%mean cambero 43 seriesx 44 ••+ 45

6Z mean

04 6 /2 16 20 4 8 /2 16 20 4 8 e 16 20 24Maximum thickness in percent of chord

FmURE 85.—Variation of maximum lift with thickness.

slope slightly. Table II gives the numerical values ofthe slope in convenient form for noting the general

trends with respect to variations in thickness and in

camber.It will be noted that all values of the slopelie below the approximate theoretical value for thin

wings, 2a per radian; the measured values lie between95 and 81 percent, approximately, of the theoretical.

The angle of zero lift is best analyzed by means of acomparison with that predicted by the theory. Thin-airfoil theory states that the angle of zero lift is pro-

portional to the camber if the camber is varied, aswith these related airfoils, by scaling the ordinates of a

perimental values lie between 100 and 75 percent,approximately, of the theoretical values, the depar-ture becoming greater with a rearward movement ofthe position of the camber and with increased thickness(above 9 to 12 percent of the chord). Numericalvalues of the angle of zero lift are given in table III.

Maximum lift.—The variation of the maximum liftcoefficient with thickness is shown in figure 85. It

will be noted that the highest values are obtained withmoderately thick sections (9 to 12 percent of the chordthick, except for the symmetrical sections for which

the highest values are obtained with somewhat thicker

s

48 REPORT NATIONAL ADVISORY ;OMMITTEE FOR AERONAUTICS

sections). The variation with camber, shown infigure 86, confirms the expected increase in maximumlift with camber. The gain is small, however, for thenormal positions of the camber, but becomes largeras the camber moves either rearward or forward. Itwill be seen by reference to figure 85 that the camberbecomes less effective as the thickness is increased.

This reduced effectiveness of the camber is in agree-

ment with a conclusion reached in reference 13 that

for airfoils having a thickness ratio of approximately20 percent of the chord, camber is of questionable

value. Numerical values of the maximum lift co-efficient are given in table IV.

Air-flow discontinuities.—These and other wind-tunnel tests indicate that at the attitude of maximum

bar (for airfoils having normal camber positions;i.e., 0.3c to 0.5c).

MOMENT

Thin-airfoil theory separates the air forces acting onany airfoil into two parts: First, the forces that pro-duce a couple but no lift (they are dependent only onthe shape of the mean line); second, the forces thatproduce the lift only, the resultant of which acts at

Maximum thickness in per cent •,-.05 i

69-.04 j.Theoretico/ curvy ' L5

w -.03 2/

^EUo -.02

-.0/

Ca„

_L4

Comber position to rrau/on or enora(Abscissa of maximum mean-line ordinate)

FIGURE 88.—Variation of maximum lift with camber. Results for 12 percent thickairfoils.

lift the air forces on certain airfoils exhibit suddenchanges which in many instances result in a serious

loss of lift. The probable cause of these air-flow dis-continuities is discussed briefly in reference 13. The

stability or instability of the air flow at maximumlift may be judged by the character of the lift-curve

peaks indicated for the various airfoils. The curves

are classified into three general types as noted intable IV, but the degree of stability is difficult tojudge. It may be generally concluded that improvedstability may be obtained by (1) having a smallleading-edge radius, which causes an early break-

down of the flow with a consequent low value of themaximum lift, (2) increasing the thickness (beyond thenormal thickness ratios), or (3) increasing the cam-

_L4

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 for12 percent thick airfoils. Craves Indicate general trends for the different thick-nasses.

.90

o ,.BL

E m

Qt

V$ C• 74

4 8 12. 16 20 a4Maximum thickness in percent of chord

FIGURE 88.—Variation of moment at zero lift with thickness. Numbers refer tomean-camber designation.

a fixed point. We then have in the working range anexpression for the total moment taken about anypoint

Cm = C,ea+nCL

where Co is the moment coefficient at zero lift andnCL is the additional moment due to lift.

As with the angle of zero lift, the theory states thatthe moment at zero lift is proportional to the camberand predicts an increase in the magnitude of themoment as the camber moves back along the chord.Figures 87 and 88 show the values of the moment

coefficient as affected by variations of camber andthickness compared with the theoretical values. Re-

ferring to figure 87, the plotted data indicate that the

0 23 —024 _1o25

0 430 44o 45

0 630 640 65

i

0 Ox mean combe .x 2%a 4%+ 6%

Symmetric./ oirfoixe * q

+ Norma/cambered oirfoi/s

.04

.02 C&-

+0.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 49

moment coefficients are nearly proportional to thecamber. It will also be noted that the curves repre-senting the ratios of the experimental coefficients tothe camber are nearly parallel to the equivalent curverepresenting the theoretical ratios except that thecurves tend to diverge for positions of the camberwell back. Figure 88 shows that the experimentalvalues lie between 87 and 64 percent, approximately,of the theoretical. Numerical values of the momentcoefficient at zero lift are given in table V.

the profile drag as the minimum value plus an addi-tional drag dependent upon the attitude of the airfoil,we have in coefficient form

CD = CD,+ (CDamtn + ACDa)

The induced-drag coefficient C U, which is computedby means of the formula given in reference 8, is con-sidered to be independent of the airfoil section. Thevariation of the profile-drag coefficient with the shapevariables of the airfoil section is analyzed with respect

4 8 0 16 20 24Maximum thickness in per cent of chord

FIGURE 89.—Variation of position of constant moment with thickness. Values ofa for equation .,,, .s+nC . Results for airfoil. having normal camberpositions (0.3c to 0.6c).

If the resultant of the lift forces acted exactlythrough the quarter-chord point, as predicted by thetheory of thin airfoils, there would be no additionalmoment due to the lift when the moments are takenabout this point. The curves of C ,, against CL,however, show a slope in the working range whichindicates that the axis of constant moment is displacedsomewhat from the quarter-chord point. The factor nrepresents the amount of this displacement as obtainedfrom the deduced slopes of the moment curves in the

Mean camber4

s%)2

Symmetrical airfoil 4 %

zr,O .2 4 .6 .8 /.O

Comber position in fraction of chord(Abscissa of maximum mean-line ordinate)

names 90.—Variation of position of constant moment with camber. Values of nfmnquationC..l,—C.,+nCz. Results for 12 percent thick airfoils.

normal working range. The variation of this dis-placement with thickness and with camber is shown infigures 89 and 90. Table VI gives the numericalvalues. Beyond the stall all the airfoils show a sharpincrease in the magnitude of the pitching moment.The suddenness of this increase follows the degree ofstability' at the stall as indicated by the type of thelift-curve peak.

DRAG

The total drag of an airfoil is considered as made upof the induced drag and the profile drag. Considering

in fractionFIGURE 91.—Varlation of minimum profile dmg with thioknom far the symmetrical

airfoils.

to the variations of the two components of the profiledrag.

Minimum profile drag.—The variation of the mini-mum profile-drag coefficient with thickness for thesymmetrical sections is shown in figure 91. The cam-bered sections show the same general variation withthickness but, to avoid-confusion, the results are notplotted. The variation of the minimum profile-drag

.00&Mean ca, bar,

V

.Co a4X

2%

O .2 .4 s .8 LLComber position in fraction of chord

(Abscissa of mu. mean-11ne ordinate)FIGURE 92.—Incmase in minimum profiledmg due to camber. Reseltsforl2pereent

thick airfoils. Values of k for equation Cx,... k+0.0060+0.01 9+01 V, where k

is the increase in CDs.., due to camber and t is the maximum thickness in fractionof chord.

coefficient with the profile thickness may be expressedby the empirical relation

CDamfn =k +0.0056+0.01t +0.112

where t is the thickness ratio and k (which is approxi-inately constant for sections having the same meanLine) represents the increase in C,,,.,. above thatcomputed for the symmetrical section of correspondingthickness. The variation of CDanfa with camber isindicated by the variation of k as shown in figure 92.

50 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

The effect of camber is small except for the highlycambered sections having the maximum camber wellback. Numerical values of CDamr° are given in table VII.

Additional proflle drag.—The additional profiledrag, which is dependent upon the attitude of the air-foil, has previously been expressed as a function of thelift (reference 4) by the equation

ACDe = CDO — CDOmtn = 0.0062 (CL — C,,.,,)'

where CL°n, may be called the optimum lift coefficient;that is, the lift coefficient corresponding to the mini-mum profile-drag coefficient. This equation holds

This function is represented in figure 93 as the curvedetermined from the results for the symmetrical air-foils and for the airfoils having a camber of 2 percentof the chord. As the camber is increased, the dis-persion of the plotted points from the curve becomesgreater. In general the points above the curve corre-spond to thick sections and sections in which the maxi-mum camber is well back. The departure from thecurve becomes greater with increased thickness andwith a rearward movement of the maximum-camberposition. The points well below the curve correspondto the thin airfoils.

.028

.024

.020

016J11.012

.008

.004

x0 OX mean

24 % .,6%

comber I ° °-1

xn

—

xX#•

:° tx

.2 .3 .4 .5 .6 .7 .8 .9 40a-a-,

F—E 93.—Addaioml pmdle drag.

reasonably well for the normally shaped airfoils atvalues of the lift coefficient below unity.

A convenient practical method of allowing for theincreased values of CD, at moderately high values ofthe lift coefficient is to include the additional profiledrag with the induced drag, as suggested in reference2. For the symmetrical airfoils of moderate thicknessthe term to be added to the induced-drag coefficientwas given as 0.0062 CLE. The relative importance ofthis term may be better appreciated by consideringthat it represents 11.7 percent of the induced drag ofan elliptical airfoil of aspect ratio 6. The samemethod may also be applied to other airfoils if thevalue of the optimum lift is not too large.

Andrews (reference 14), using the part of these datapublished in references 2, 4, and 5, suggests for theadditional profile drag the form

A D0 =f1 CL— CLODr

CL _%L,

1

Because the additional profile drag is not a simplefunction of the lift, and also because the results aspresented in figure 93 are difficult to follow, generalizedcurves for the relation

OCDe = f (CL— CLavr)

are given in figure 94. These curves are given torepresent more accurately the additional profile dragfor the normally shaped sections.

Optimum lift.—The optimum lift, as defined above,is the value of the lift corresponding to the minimumprofile drag. As the determination of this value of thelift is largely dependent upon the fairing of the profile-drag curves, special curves were faired for this purposeon enlarged-scale plots corresponding to certain relatedairfoils grouped together. The values of the optimumlift coefficients obtained in this manner are given intable VIII. It may be noted by reference to this tablethat the optimum lift coefficient increases with camber

CHARACTERISTICS OF AIRFOIL SECTIONS FRO]

and for the highly cambered sections a definite increaseaccompanies a forward movement of the camber.

.02

.0/

0.02

.o/

0.02

.0/

VgI 0

x$.02

.0/

O

.02

.0/

0

.02

.Ol

O 2 .4 .6 .8 10 /.2 14

FIGURE 91.—Additional profile dmg as a function of Cy— C.,1. Results era for sit-falls having normal camber positions (0.5c to 0.5c).

I TESTS IN VARIABLE-DENSITY WIND TUNNEL 51

it is not primarily dependent upon the shape of themean line. Nevertheless it is interesting to comparethe optimum lift coefficients with the values includedin table VIII representing the theoretical lift coeffi-

cients at the "ideal" angle of attack for the mean

line; i.e., the angle of attack for which the thin-airfoiltheory gives a finite velocity at the nose. (See the

appendix.)

GENERAL EFFICIENCY

The general efficiency of an airfoil cannot be ex-pressed by means of a single number. The ratio of

Symmetrical airfoil

^r

1

Meon rnmber

1

i

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 pereeutthick airfoils.

the maximum lift to the minimum profile drag is, how-ever, of some value as the measure of the efficiency ofan airfoil section. The variation of this ratio withthickness is shown in figure 95. The curves of this

figure indicate that the highest values of the ratio are

2..16.v-mapu

"eon0 07. comberx 29,—••

I.

o 49,+ 6%—IT

Max, lhickness 69. 4.

,*

I "-T1 1 1 1 6 410'2-

Maximum lhickness 9%0

`6 4° 2 0°

Maximum lhickness /2% k

E

6 4 2 f)-

^^

Maximum lhickness /5%ov^F

a.rA `. er• UU

6 O,

Maximum thickness /B9,o w

$ ^o

.oFmo.

u

1 4° 2 O•.

Maximum thickness 219,O

20

G°

J B

4

200

160

V^/20

v 80

40

Symmetricalseries

+25 4 +

2%mean comber 4%mean comber 6%mean comber0 23 series o 43 series o 63 seriesx 24 x 44 x 64

_j

0

+65

4 8 12 16 20 4 6 /2 /6 20 4 8 12 16 20 24Maximum 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 thevariation with thickness. The rapid decrease in the chord thick. The variation with camber, shown inoptimum lift with increased thickness indicates that figure 96, is less important. An increase in the camber

52 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)tend to decrease the value of CL,aee/CDO,afa. Thenumerical values of the ratio are given in table IX.

SUPPLEMENTARY AIRFOILS

For the purpose of investigating briefly the effects

of certain shape variables other than those discussed

in the main body of the report, 10 supplementary air-

foils were tested. The airfoil sections were as follows:6 symmetrical sections with modified nose shapes, 2sections with reflexed mean lines, and 2 sections simu-lating those of a wing having a flexible trailing edge.

Maximum thickness 0.12c

/.4^Moximum thickness 0.18c

/.2

/.0

Moximum thickness 006e

.6

.4

.2

O / 2 3 4 5 6 7i-Hnn-eons radus in Der cent of chord

F-RE 97.-Variation of maximum lift with nose radius.

Airfoils with modified nose shapes.-The airfoils ofthe first supplementary group investigated were de-veloped from three of the symmetrical N.A.C.A. familyairfoils: The N.A.C.A. 0006, the N.A.C.A. 0012, andthe N.A.C.A. 0018. For each of these basic (or nor-

mal) sections one thinner-nosed section, denoted by

the suffix T, and one blunter-nosed section, denoted

by the suffix B, were developed and tested. Thederivation of each modified section was similar to thatof the normal section and was accomplished by a sys-tematic change in the equation that defines the normalsection. This change is principally a change in thenose radius, but it also results in modifications to theprofile throughout its length, except at the maximumordinate and at the trailing edge. The nose radii of thesections in percent of the chord are as follows:

seetlon T series Normal 1111-11

0006 0.10 0.40 1.1900120018

.40

.891.583.56

3.807.15

The aerodynamic characteristics of the modifiedsections are given in figures 72 to 77. These may becompared with the characteristics of the normal sec-tions given in figures 4, 6, and 8. The maximum liftcoefficients of the modified and the normal sectionsare plotted against the leading-edge rad ii in figure 97.

It is interesting to note that the leading-edge radius isvery critical in its effect on the maximum lift whenthe radius is small. This critical effect is also indicatedby the rapid increase in the maximum lift with increas-ing thickness for the thin sections as shown in figure 85.

Airfoils with reflexed mean lines.-Previous inves-tigations have shown that the pitching moment ofcambered airfoils can be reduced by altering the formof the mean line toward the trailing edge, with a con-sequent loss of maximum lift but only a small reductionin 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 thebasic thickness distribution of the N.A.C.A. 0012 dis-posed about certain mean lines of the form given inreference 15

y^=hx(1-x) (1-Tx)

The values of h in this equation were chosen to give acamber of 0.02 and the values of X were chosen to givethe airfoil designated the N.A.C.A. 2R 1 12 a smallnegative moment and the airfoil designated theN.A.C.A. 2R 212 a small positive moment. Charac-teristic curves for the two airfoils are given in figures78 and 79. The principal characteristics of the sec-tions may be conveniently compared with those of therelated symmetrical section, the N.A.C.A. 0012, anda related normal section having a camber of 2 percent

of the chord, the N.A.C.A. 2412, by means of the

following table arranged in the order of increasing

pitching-moment coefficients.

c.section Cam„ CDC- CDC- C.p

211M 1.47 0.0086 171 0.00400122R^12

1.531.63

.0083

.0083184184

-.002-.020

2912 1.02 .0085 IN -.044

These results indicate that airfoils having reflexed

mean lines may be of questionable value because of theadverse effect of this mean-line shape on the maximumlift coefficient.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 53

Thickness and camber modifications near thetrailing edge.—Two airfoils were developed to simu-late an airfoil having a flexible trailing edge in a straightand in a given deflected position. The thickness dis-tribution is composed of three parts: the forward por-tion (0 to 0.3c) having the same distribution as theN.A.C.A. 0012, the rear portion (from 0.7c to thetrailing edge) having a thin, uniform value, and thecentral portion joining these two with fair curves.As shown in figure 80, the two airfoils differ only inthe rear portion, the section designated N.A.C.A.0012Fo simulating that of a wing having the trailingedge deformed for the high-speed condition, and thesection designated N.A.C.A. 0012F I simulating thatof the same wing with the trailing edge bent down in acircular arc. Curves of the aerodynamic characteris-tics for both conditions are compared in figure 80.Considering the results given by both airfoils as twoconditions for one airfoil, a very high maximum liftwith a reasonably low minimum drag is obtained.

On this assumption the ratio CL'-C."ef. Is

197, slightly

higher than the value of this ratio given by theN.A.C.A. 2412.

In order to study the effects of an extreme change inthe thickness distribution, the principal characteris-tics of the two sections may be compared with thoseof the related normal sections, the N.A.C.A. 0012 andthe N.A.C.A. 6712. The maximum lift coefficient islittle affected by the change in the thickness distribu-tion, but it is of interest to note (table I) that the slopeof the lift curve of the N.A.C.A. 0012Fa is slightlygreater than 2a per radian, as compared with an appre-ciably lower slope for the N.A.C.A. 0012. The profiledrag is also affected by the change in the thicknessdistribution. Of the two symmetrical sections, theprofile drag of the N.A.C.A. 0012Fp is much higherthan that of the N.A.C.A. 0012 over the entire liftrange. This is not true, however, for the two cam-bered sections. Comparing the characteristics of theN.A.C.A. 0012F, with those of the N.A.C.A. 6712, wefind that at low values of the lift the profile drag ofthe former is much higher, but as the lift increases thisdifference becomes less, and in the high-lift range theprofile drag of the N.A.C.A. 0012F, is considerablyless than that of the N.A.C.A. 6712.

CONCLUSIONS

The variation of the aerodynamic characteristics ofthe related airfoils with the geometric characteristicsinvestigated may be summarized as follows:

Variation with thickness ratio:1. The slope of the lift curve in the normal working

range decreases with increased thickness, varyingfrom 95 to 81 percent, approximately, of the theoreticalslope for thin airfoils (2v per radian).

2. The angle of zero lift moves toward zero withincreased thickness (above 9 to 12 percent of thechord thickness ratios).

3. The highest values of the maximum lift are ob-tained with sections of normal thickness ratios (9 to15 percent).

4. The greatest instability of the air flow at maxi-mum lift is encountered with the moderately thick,low-cambered sections.

5. The magnitude of the moment at zero lift de-creases with increased thickness, varying from 87 to64 percent, approximately (for normally shaped air-foils), of the values obtained by thin-airfoil theory.

6. The axis of constant moment usually passesslightly forward of the quarter-chord point, the dis-placement increasing with increased thickness.

7. The minimum profile drag varies with thicknessapproximately in accordance with the expression

C- i.,n =k+0.0056 +0.0lt+0.lt2

where the value of k depends upon'the camber and t isthe ratio of the maximum thickness to the chord.

8. The optimum lift coefficient (the lift coefficientcorresponding to the minimum profile-drag coefficient)approaches zero as the thickness is increased.

9. The ratio of the maximum lift to the minimumprofile drag is highest for airfoils of medium thicknessratios (9 to 12 percent).

Variation with camber:1. The slope of the lift curve in the normal working

range is little affected by the camber; a slight decreasein the slope is indicated as the position of the cambermoves back.

2. The angle of zero lift is between 100 and 75 per-cent, approximately, of the value given by thin-air-foil theory, the smaller departures being for airfoilswith the normal camber positions.

3. The maximum lift increases with increased cam-ber, the increase being more rapid as the cambermoves forward or back from a point near the 0.3cposition.

4. Greater stability of the air flow at maximum liftis obtained with increased camber if the camber is inthe normal positions (0.3c to 0.5c).

5. The moment at zero lift is nearly proportional tothe camber. For any given thickness, the differencebetween the experimental value of the constant ofproportionality and the value predicted by thin-airfoiltheory is not appreciably affected by the position ofthe camber except for the sections having the maximumcamber well back, where the difference becomesslightly greater.

6. The axis of constant moment moves forward asthe camber moves back.

7. The minimum profile drag increases with in-creased camber, and also with a rearward movementof the camber.

54 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

8. The optimum lift cofficient increases with thecamber and for the highly cambered sections a definiteincrease accompanies a forward movement of thecamber.

9. The ratio of the maximum lift to the minimumprofile drag tends to decrease with increased camber(above 2 percent of the chord) and with a rearwardmovement of the camber (for the highly camberedsections).

LANGLEY MEMORIAL AERONAUTICAL LABRATORY,NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS,

LANGLEY FIELD, VA., December 20, 1932.

APPENDIX

It is proposed in this section of thereport to present,briefly, a summary of the results of the existing thin-airfoil theory (based on the section mean line) as ap-plied to the prediction of certain section characteris-tics. Such a summary is desirable because at presentthe results must be obtained from several differentsources which give them in a form not easily applied.Three characteristics are considered; namely, (1) theangle of zero lift aLO, (2) the pitching-moment co-efficient C d4, and (3) the "ideal" angle of attack a,,or the corresponding lift coefficient Czl, that is, valuescorresponding to the unique condition for which thetheory gives a finite velocity at the nose of the airfoil.(See reference 16.)

Expressions for lift and moment coefficients may bewritten as follows if the angles are measured in radians:

Cy=27r(a-%) (1)

CLl=27r(aj-a4) (2)

Gme/4 = 2(13+aL0) (3)

If the leading end of the mean line is chosen as theorigin of coordinates and the trailing end is taken onthe xaxis at x=1, then the parameters aLo, a,, andj3 are given by the following integrals

azo = fly fl (x) dx0

(4)

ar = f y f2(x) dx0

(5)

d= ft yf3(x) dx (6)

0

where-1

fl(x)= 7r(1-x) [x(1-x)]E (7)

(1- 2x)f2(x)=2,'[x(1-x)]n (8)

f3(x)- 4(1-2x)

(9)7[x(1-x)]11

and y is the ordinate of the mean line at a given abscissax. The integrals (4) and (6) may be shown to beidentical with the corresponding integrals given byGlauert (reference 15) and by Munk (reference 17),and integral (5) is given by Theodorsen (reference 16).

The evaluation of these integrals for the N.A.C.A.airfoil sections given in this report was accomplishedanalytically. The values of aL, (changed fromradians to degrees), C-c4 and CL„ so computed, aregiven in tables III, V, and VIII, respectively, in themain body of the report. This method of evaluation,however, cannot be applied to many of the commonlyused sections because they do not have analyticallydefined mean lines; hence, an approximate methodmust be used. A graphical determination gives goodresults and for convenience the values of the threefunctions, (7), (8), and (9), at several values of x, aregiven in the following table:

x f1W r,(x) f3W x f,(-) W4 W4

0 0.30 -0.992 0.662 1.1110.0125.0260 -2.901-2.091 118.1539.73 111.177.747 .40.60 -1.083-1.273 .2710 .5200.0500.0750 -1.537-1.308 13.847.403 5.2589.109 .60.70 -1.624-2.315 -.271-.662 -.620-1.111

.1000 -1.178 4.716 3.395 .80 -3.978 -1.492 -1.910.15.20 -1.0482.995 2.4471.492 2.4961.910 .90.95 -10.61-29.21 -4.718-13.84

-3.395-5.258.25 -.9so .980 1.470 1.00

In general, some difficulty would be expected withthe graphical method because the values of the abovefunctions tend to infinity at the leading and trailingedges. Actually, because the ordinates of the mean-line extremities are zero, the integrand may approachzero, and does at the leading edge for the integral (4),and at the leading and trailing edges for the integral (6).Difficulty, however, is encountered at the trailing edgefor the integral (4) and at the leading and trailing edgesfor the integral (5). In order to avoid this difficulty,integral (4) is evaluated graphically from x=0 tox=0.95, and the increment contributed by the por-tion from x=0.95 to x=1 is determined analytically.Likewise, integral (5) is evaluated graphically fromx=0.05 to x=0.95 and analytically for the extremities.The analytical determination of the increments isaccomplished by assuming the mean line near theends to be of the form

y=a+bx+cc2

Evaluating the integrals gives

daze = -0.964yo .s5 +0.0954y; (x=0.95 to x=1)

Aa,= +0.467y,.05 +0.0472yo (x=0 to x=0.05)

I -0.467yo.95 +0.0472y1 (x=0.95 to x=1)

where yo and yf are the mean-line slopes at theleading and trailing edges, respectively.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TESTS IN VARIABLE-DENSITY WIND TUNNEL 55

REFERENCES

1. Jacobs, Eastman N., and Anderson, Raymond F.: Large-Scale Aerodynamic Characteristics of Airfoils as Testedin the Variable-Density Wind Tunnel. T.R. No. 352,N.A.C.A., 1930.

2. Jacobs, Eastman N.: Tests of Six Symmetrical Airfoils inthe Variable-Density Wind Tunnel. T.N. No. 385,N.A.C.A., 1931.

3. Pinkerton, Robert M.: Effect of Nose Shape on the Char-acteristics of Symmetrical Airfoils. T.N. No. 386,N.A.C.A., 1931.

4. Jacobs, Eastman N., and Pinkerton, Robert M.: Tests ofN.A.C.A. Airfoils in the Variable-Density Wind Tunnel.Series 43 and 63. T.N. No. 391, N.A.C.A., 1931.

5. Jacobs, Eastman N., and Pinkerton, Robert M.: Tests ofN.A.C.A. Airfoils in the Variable-Density Wind Tunnel.Series 45 and 65. T.N. No. 392, N.A.C.A., 1931.

6. Jacobs, Eastman N., and Pinkerton, Robert M.: Tests ofN.A.C.A. Airfoils in the Variable-Density Wind Tunnel.Series 44 and 64. T.N. No. 401, N.A.C.A., 1931.

7. Jacobs, Eastman N., and Ward, Kenneth E.: Tests ofN.A.C.A. Airfoils in the Variable-Density Wind Tunnel.Series 24. T.N. No. 404, N.A.C.A., 1932.

S. Jacobs, Eastman N., and Abbott, Ira H.: The N.A.C.A.Variable-Density Wind Tunnel. T.R. No.416, N.A.C.A.,1932.

9. Higgins, George J.: The Prediction of Airfoil Characteris-tics. T.R. No. 312, N.A.C.A., 1929.

10. Knight, Montgomery, and Harris, Thomas A.: Experi-mental Determination of Jet Boundary,Corrections forAirfoil Tests in Four Open Wind Tunnel Jets of DifferentShapes. T.R. No. 361, N.A.C.A., 1930.

11. Stack, John: Tests in the Variable-Density Wind Tunnel toInvestigate the Effects of Scale and Turbulence on Air-foil Characteristics. T.N. No. 364, N.A.C.A., 1931.

12. Dryden, H. L., and Kuethe, A. M.: Effect of Turbulence inWind Tunnel Measurements. T.R. No. 342, N.A.C.A.,1930.

13. Jacobs, Eastman N.: The Aerodynamic Characteristics ofEight Very Thick Airfoils from Tests in the Variable-Density Wind Tunnel. T.R. No. 391, N.A.C.A., 1931.

14. Andrews, W. R.: The Estimation of Profile Drag. Flight,vol. XXIV, no. 25, pp. 530a-530d, 1932 and no. 31, pp.710a-710c, 1932.

15. Glauert, H.: The Elements of Aerofoil and Aircrew Theory.Cambridge University Press (London), 1926.

16. Theodorsen, Theodore: On the Theory of Wing Sectionswith Particular Reference to the Lift Distribution.T.R. No. 383, N.A.C.A., 1931.

17. Munk, Max M.: Elements of the Wing Section Theory andof the Wing Theorv. T.R. No. 191, N.A.C.A., 1924.

E. MI i

HIR A

c,

?a r-.r -.-.8 v

eva 2 a 8 t aa

cp

....

....

....

...

sees

...........

....

....

....

....

115 s

tngR

ttSS2 22

tgma

-8

I-ri

Irri

II

II

IIrl

II

IIrl

II

II

II

II

II

I

.....

.^cl

- 9 . 2

. 2 - S

. 2 - 2

-S - 2 -S

- S! - 2

- 2 - S

-S

- !2 -2-2.2

-2 - e

- 2

-S - 9 ......; .....................

2

- -----

. .... .. .

.. ....

4aa3

^5tl

2 i

^k^t

-,st

al

N52

Rns4

:- . 2 age

t3 ^3

iR iR vt

S a v

tgu

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 2

vM

222

H M

99

99

2i 9

!-f

E i

Rt.2

2.

CHARACTERISTICS OF AIRFOIL SECTIONS FROM TEST$ IN VARIABLE-DENSITY WIND TUNNEL 57

TABLE IL-SLOPE OF LIFT CURVE, ae= dC^ (PER DEG.)dao

Thicknessdesigna-

tgnntinn

06 gg 12 15 Is 21 25 112

011 -___________ 0.102 0.101 0.101 0.100 0.098 0.094 0.089 0.101

22- ------- -------

.10323::::::::::::

25:f031

03

--.-l0,3--

02

-".-1,0,2--l 02

7-10-2-:lot099 ON

:097--------

------

.101110102

26 ____________ ------- ------- --om.

0027____________ ------- ------- -- 00

42------------ ------- ------- ------- ------- - ------- --43---- -------

102 183 10% : 1,21 'm : 0,03,I :::-, 'OO'D:

. 1,3044------------46____________

:.104

:.103 .101 .096 .095 :017

047____________ ----- --------------

-

------------------- -

- ---------- -------

098.097

82.....-_____-------- - -- ----- --- --- - -------1

84 l04.104.101

H6.102:1. 7

OD9ig

.099.096: OD6

:10I: 101

.101 .103 .101 . 099 .095 094 101OK

1.67------------ --- - -- --- - --- 097

I Additional tests to determine variation with camber.

TABLE III.-ANGLE OF ZERO LIFT, .4 (DEGREES)

Thickness

desire-\ t on

b ell desX

ig-tion

06 09 12 15 18 21 25 12 Theor.

00___-__._.___ -0.1 0.0 0.0 0.0 0.0 -0.1 0.0 0.0 0

22---- - - - ---- ------------- -------- ---------- ----- '

- go

-1.23.__________-24. __________-

-1.8-1.7

-2. 0-1.7

--1.7 -1.7 -1.9 -1.7 ------

9-1.8 -2 08

25_ -____.___-_26 -----------

-20-------

-2.0-------

-2. 0-

-2.0 -20 -1.8 ------ -2. 13

-2.29 2. 59

27 ------------ ------- ------- - ::::: :::-:: :::::: :.. -N -3.04

42------------ ------- ------- ------ ------ ------ ------ ------ -3.4 -3.6043 -3.8 -3.6 -3.7 -3.6 -3.5 -3.6 ------ 3 -3.84

46......... ._. -3. 9 4.

3-3.- 4.1

8.1- 4.

1. 8

-4.1-1_,: 9 -3. 4

-3. 4---... -4.2

.15-4.58

40_ ___________ ------- ------- - -4.6 -5.1847____________ ------- ------- - -5.0 -6.09

92 ----------- ------- ------- ------ ------ ------ ---------- -6.2 -5.40-b.2 -5.4 -5.4 -5.4 --8.2 -6.2 ---- -5.5 -5.76

1466...._._.._.

-1:-8.3 6.

6 -1.1_3

-6.1_6.

-^I: 70

-1-7-5.7

-I 2-b.3

-1-72

-6.23-6.88

66.__________ ------- ------- -- 6 -7.7887_ ___________ ------- -------

-- :::- :::-::7.0 -9.13

I Based on straight portion of lift curve extended. Bee curve for actual value.

TABLE IV.-MAXIMUM LIFT COEFFICIENT, CL..

d\Th'Cbb d__or des.

- - 13 M A , Z 12

00 .---..-.._._ • 0.98 x 1.27 -1.53 1 1.53 x 1.49 - Laii 1 1.2() -1.53

22 ___ _________ ------- ------- ------- - ----- - - --- ------- ------- 11.60

23 ..-______.__ • 1.04 -1.61 11.60 °1.54 ------- ------- -'1:24_____.._----

25 -----------• 1.01

• 1.03

11.51

-1.38

1 1.59

1 1.60

1 1.55

1 1.53

1 1.43• 1.48

-1.36

1 1.38

-::::-:

-------

1 211.02

26- -----------27------ ----

-------

------- ---------------------

---------------------- ------

--------------

--------- :::::: 1: 8:168

42____________ ------- ------- ------ ------- ------- ------- ------- • 1.71

43_______..._. -I :M 160 b1: 63 lf : 51 11-41 1:1:l 11

4445

1

:1160

•1

1 :

. 1.56

^1.61. 1.69

17

62• 1.47. 1

A 417

•1.4

6 ------11

1 1.69

46 ............ .... -- -----

1171

47____________---- .. .....

• 1. 82

62_ ___________ ------- ------- ------- ------- ------- ------- ------- 1 1.75

b411 :

43

11.87 1l.64 ^l.65 1: 43 1 1.37 11: 116465___._ _ ___.__ -1.29

11

1 1: 6781 •1:

85

75 16-1:897

:, ,

-1.61

: 1- 41•1.49

:::::::

-------<I.

61 1.75

66 --- ------- ------- ------- ------- ------- ------- ------- - - 1.83

67 ............ ............ . ...... . ........ .............. .

95

NOTE-I,ettff indicates type of lift enrva, peak.

I Theoretical lift coefficient at "Ideal" ougle of attack.

TABLE IX.-RATIO OF MAXIMUM LIFT COEFFICIENTTO MINIMUM PROFILE-DRAG COEFFICIENT,CL... 147DO.-

Thicknessd

I \^d

gs gg igis

18

S.

26 12

00-_-____...__ 185 172 184 164 138 115 84 184

22------------ ------- ------- -- 1842a ------------ 142 182 18124_____.___ 144 189 177 156 128 106 ------- 10026 14 170 180 148 132 109 ------- 18420___ _ ________

_ -_ __

27____________ ------- ------- ------- ------- ------- ------- ------- 187---- ----- - _ ----- - _ ----- ------- ------- ------- 187

42____________ ------- ------- - ----- ---- -- ------- ------- ------- 18643 --__-_______ 161 140 123 96 ------- 17244______-- :::: IN 1,82 170 150 127 104 ------- 17946. .______.___ 132 169 164 143 123 106 ------- 17846------------ ------- -------------- ------- --------------------- 1784>____________ ------- ------- ------- ------- ------- ------- ------- 175

82.------- ---- __ __i6__ ­iii­ _' iiF ii6 ii_ ____ ___---------

'a

63- -- -iii____- -

I-64.-___--_._-. 160 179 159 Isla 114

1`7 : -----

I.65..__.....__. lag 171 158 132 114 97

58 REPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

TABLID V.-MOMENT COEFFICIENT AT ZERO LIFT, C.,

Thickness

ber designation

06 09 12 16 18 21 25 12 Theor.

00_. ._.____. -0.002 -0. ON -0.002 0.000 -0. 002 -0.001 -0.003 -0.002 0

22------------ ------------ ------------ ------------ ------------ ------------ ------------ ------------ -0.029 -0.037023 .__.._.____. -.036 -.038 0136 -.034 ------------ ----- - ------ ----- - ------ -.038 - . 0447 24__._-______25 -- - --- - ----

-.039-.048

-.044­ 062

W-.050

-.010-.019

-.037 -.047

-.036-.043

-- - ---------------------

-.044

-.054 -.0531

om28_ ___________ ____________ ---- -------- ------------ ------------ ------------ ------------ ------------ -.000 074927____________ ------------ ------------ ------------ - - - --------- ------------ ---- -------- ------------ - . 076 -.0912

41____________ ------------ ----- ------ ------------ --- --------- ------- ----- ------- ----- ------------ -.059 -.073943 .___.__.____ -.075 -.073 -.072 -.008 -.066 -.057 -------- ---- -.075 -.099444------------46_-._.__-___

-.087- 109

- 086-:106

- . 087-:102

-.083-.097

-:078- . 094

-.071-:082

---- : --- : ------- --- - - -

-.039-:105

-.1062-.1257

46 ____________ ____________------------ ------------ ------------ ------------ ------------ -------- ---

-.124 -.149747____________

------------ --- - -------- ---- --- - ---- ----- - ------ ------------ -------------------- -----.143 -.1825

62____________------- - ---- --------- ---

------------ ----- - ------ ------- ----- ------------------------ _.087 -.110963_____.__-__. 64. . ____.___._ III 29

110Ias

il29

-:106-.125

-:097-.118

-.090-.110

------------.... -------

_.II0-.132

-.1342-.1594

65_______ __._ 159 .158 -.154 x-.160 -.139 -.129 __ - . 16986_.________67 ------ ----------

:186-.206

2W2737

A Based on straight portion 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 FOREQUATION Cm04 =

C,+nCL) Td.Thk__designs

Thick-Thickness

designs-if..

Coca

•ber des-fgnation

ionis

a - ed

Ig

00_--_.____-_- 0. 7 0.7 0.9 -1.1 1.4 IT - -0.9

22------ --23::::::::::::3 ------ --3 ------ --3 ------5--

24____________25 .-_______ ._

.1 .0

.3

.2.4.3

.7

.61.01.0

1.51.7

.......-------

.0

.726 ----------- ------- ------- ------- ------- ------- .827 ----- ----- ------- ----- - ------- ------- ------- .8

42 ------ ---- ------- ------- ------- ------- ------43____________44____________

.3.3

.4

.3.55

.710

1.2L4

1.61,7 :5

4646_ __________

:: -------- -1 .3-----

: - 8 -------

:.

9 1.4 1.71.0

'47___________ ------- ----- I:

62 ------ ------- . 2

64____________ ­ 7 .0 .6 .9 1.3 1.7 ------- .886____________66__ ___ __ _

.0 .0 .7 1.6 1.8 1.9

----

------- 1.5

67:-:.::.. 1^^ -------- --------------- -------- 1:::::::Il of2:

TABLEV II.-MINIMUM PROFILE-DRAG COEFFI-CIENT, Cna_

igi

C

_ '1" 06 09 12 15 18 21 25 12

Thicknessdes na

bar des,Ignation,

00._.____.--__ 0..

0.

0074 0.0093 0.0093 0.0108 0.0120 0.0143 ^ 0083

22----------- ------- ------- ------- ------- --- --- ------- -------•0087

24:. 0070 0080 0090 0099 .0112 .0127 ------ .009523 :0073 : 0083 : 00ag : 0100 ------- - ----- ------- .0089

26___________ .0073 .0081 .0099 .0103 .0112 .012626 00"27____________ 0090

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