Aerodynamic Characteristics at High and Low Subsonic Mach Numbers of the NACA 0012, 64(sub 2)-015, and 64(sub 3)-018 Airfoil Sections at Angles of Attack from -2 Degrees to 30 Degrees

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RESEARCH MEMORANDUM

AERODYNAMIC JXARACTERISTICS AT HIGH AND LOW SUBSONICMACH

NUMBERS OF TEiE NACA 0012, 642-015, AND 643-018 AIRFOIL

SECTIONS AT ANGLES OF ATTACK FROM -20 'IW 300

By Chris C. Critzos

Langley Aeronautical Laboratory

Langley Field, Va.

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NATIONAL ‘ ADVISORY COMMITTEE +

FOR AERONAUTICS

WASHINGTON

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NACA RM L54H?6a

- -

NATIONAL ADYISORY COMMITTEEFOR AERONAUTICS

RESEARCHMEMORANDUM

AERODYNAMIC CHARACTERISTICS AT HIGH AND LOW SUBSONICMACH

NUMBERS OF THE NACA 0012, 64,-015, AND 643-018 AIRFOIL

-.<.-. _

SECTIONS AT ANGLES OF ATTACK FROM -2' TO 30'

By Chris C. Critzos

SUMMARY

An investigation has been made in the Langley low-turbulence pres-

sure tunnel of the aerodynamic characteristics of the NACA 0012, 64,-015, and 643-018 airfoil sections.

Data were obtained at Mach numbers from

0.3 to that for tunnel choke, at angles of attack from -2' to 30°, and

with the surface. of each airfoil smooth-and with roughness applied at

The Reynolds numbers of the tests ranged from 0.8 x 106

the leading edge.

to 4.4 x 10% The results are presented as variations of lift, drag, and

quarter-chord pitching-moment coefficients with Mach number.

INTRODUCTION

The trend of present helicopter designs toward higher forward speeds

and higher rotor-blade speeds has resulted in a need for two-dimensional

airfoil data throughout wide subsonic Mach number and angle-of-attack

ranges. In order to supply thj.s need, a number of NACA airfoil sections,

which might be used for helicopter rotors, have been investigated in the

Langley low-turbulence pressure tunnel. The results obtained with four

of these sections are reported in reference 1. The results obtained

with three additional sections, consisting of the NACA 0012, 642-015, and 64 -018 airfoil sections, are presented herein.

The aerodynamic characteristics of the three airfoil sections were

obtained at Mach numbers from 0.3 to that for tunnel choke, at angles

of attack from -2' to 30°, and with the surface of each airfoil smooth

and with roughness applied at the leading edge. The results are presented

as variations of lift, drag, and quarter-chord pitching-moment coefficients

with Mach number. In order to expedite publication of these basic data,

2 NACA RM L54HO6a

the preparation of charts having quantities other than the Mach number

‘ .- l

as the independent variable has been deferred, as has any discussion of

. . . .

the results.

.: .

. . .

. . . .

SYMBOLS

C airfoil chord

section drag coefficient

Cd

section lift coefficient

?L

section quarter-chord pitching-moment coefficient

x/4

M free-stream Mach number

R Reynolds number based on airfoil chord

a section angle of attack

APPARATUS, TESTS, AND METHOIX

The present investigation was conducted in the Langley low-turbulence

pressure tunnel with Freon-12 as the test medium. The investigation con-

sisted of measurements of the lift, drag, and quarter-chord pitching moment

of three two-dimensional airfoils at Mach numbers from 0.3 to that for

tunnel choke and at angles of attack from -2' to 30'. The two-dimensional

models consisted of the NACA 0012, 642-015, and 6h3-018 airfoil sections,

the coordinates for which are presented in table I. The models were

machined from solid aluminum alloy.

Data were obtained with the airfoil surfaces smooth and with roughness

applied at the leading edge.

For the tests with the model surfaces smooth,

the surfaces were polished to a high degree of smoothness at the time of

model installation in the tunnel. The drag coefficients measured, however,

probably do not correspond to extensive regions of lsminar flow since use

of Freon-12 as a test medium makes unfeasible the almost continuous atten-

tion to model surface condition which is required in order to maintain

extensive laminar layers. For the tests with roughness applied at the

leading edge, the roughness consisted of O.Oll-inch-diameter Carborundum

grains spread over a surface length of 8 percent of the chord back from

the leading edge on the upper and lower surfaces. The grains were thinly

spread to cover from 5 percent to 10 percent of this area.

N A C A R M L54H06a

The Reynolds numbers in the present tests ranged from 0.8 x lo6

to 4.4 x 106. The variation of Reynolds number with Mach number is

b..

shown in figure 1 for both low (-2O to 14') and high (11' to 30') angles

of attack. The difference in Reynolds numbers for the two angle-of-attack

I.. I ranges resulted from the higher stagnation pressures used in the tests at

8..

the low angles of attack.

Additional information on the testing technique is contained in

reference 1.

RESULTS

The variations of lift coefficient, drag coefficient, and quarter-

chord pitching-moment coefficient with Mach number are presented in

figures 2 to 4 for the three airfoil sections of the present investigation.

As discussed in reference 1, corrections to the data have been applied

for tunnel-wall effects and for converting the data (which were obtained

with Freon-12 as the test medium) to equivalent air results. The varia-

tions of the aerodynam ic characteristics with Mach number (figs. 2 to 4)

for some angles of attack Qere obtained from cross plots and are presented

as lines without data point symbols.

Based on the capability of the balance used to measure the lift and

drag forces and the pitching moment, the accuracies of the measurements

for various test conditions are indicated in the following table:

Accuracies of measurements

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*0.003

0.30 *0.013 +0.0030

k.003 k.0006 *.001

+.002 k.0004 +.od1

.85

As can be seen, the accuracy in the measurement of drag is rather poor

at low Mach numbers; however, at higher Mach numbers, in the region of

the force break, the accuracy of the drag measurements is within accept-

able lim its. As in reference 1, the highest Mach numbers for which data

are presented correspond to tunnel-choked conditions.

The highest Mach

number for which the data may be considered reliable is open to some

a- -

4 NACA R&i L54H06a

=’ Y

question. A Mach number 0.03 less than that for choke, at low and

. . .

) .D

. moderate angles of attack, has often been considered as a rough upper

1='

limit beyond which little confidence should be placed in the results.

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Results at high angles of attack are involved with unknown corrections

,:.E.

which are still under study.

. . .

~

. . . .

-..: Langley Aeronautical Laboratory, National Advisory Committee for Aeronautics, Langley Field, Va., July 23, 1934.

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Chris C. Critzos

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Mechanical Engineer

--4Mu

Approved: I

e- Eugene C. Drsley

Chief of Full Scale Research Division

epr

REFERENCE

1. Wilson, Homer B., Jr., and Horton, Elmer A.: Aerodynamic Character-

istics at High and Low Subsonic Mach Numbers of Four NACA 6-Series

Airfoil Sections at Angles of Attack From -2' to 31'. NACA

RM L5X20, 1953.

NACARM L54H06a

TABLE I

COORDINATES OF NACA AIRFOIL SECTIONS TESTED

(Dimensions given in percent chord)

Chordwise

Upper and mer surfact 3 ordinates

S.tation

0012 643.018

642-015

0 0

Iwe

1.208 l 5

m-m

1. 6

.75

1.25 1.894 %

ii:5228

2.615

2.5 5.0 3.555 0504

.240

z

1zm5

15 45.7”8;

6:480

7. J-9

w-w 3;

z

;*$

z

5.803 w-w

7122 t

5 0294

W-M

to%

60 :620

4.543

m-w

l. .895

3.664 4.113

m-m

3.296 Y

is 2.623

2.472 we-

8; 1.677 1.951

1. 48 1.101

8 :; ii

l 0 7 z o.400

1% .126 0

_-..-- - ..- J. E. rad.

1.580

2.208

1.590 . e.-.-- .___

..". .~.-_. .--. _ .. -. ..-~F-_I

4.8 x 106

1’

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.angle-of-attack range .-----.

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Mach number, M

Figure l.-

Variation of Reynolds number with Mach number for two angle-

of-attack ranges.

Unflagged symbols and lines denote smooth condition Flagged symbols and lines denote leading-edge roughness R = 0.8 x 106 to 2.2 x 106 -------- R = 1.6 x 106 to 4.4 x lo6 m, degrl.4 30 D 1.2 i 1.0 1.2

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L4 .6 \ .8 .‘ .3 .4 .5 .6 .7 I, Mach number, M Mach number, M f 'I (a) Section lift coefficient.

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Figure 2.- Aerodynamic.characteristics at various angles of at&ack

obtained with l.O-foot chord NACA 0012 airfoil section.

I- :' Unflagged symbols and lines denote smooth ( edition Flagged symbols and lines denote leading-y e roughness

ret

I ,a r .- .- a f ..-I n ‘ d .rl 0 .rl .3 r: z

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.02- d---_ p 0 1.-.-r- d q T.. I 1 0 g i,, I ,s $- .5 .6 . .5 .6 .7 .0 .9 Mach number, M Mach number, M (b) Section drag coefficient.

Figure 2.- Continued.

'JnflP~gsd symbdls and liner denote smoth condition Plqged symbols and lima denote lsadlng-adga roughne,a.

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R = 0.8 x 106 to 2.2 ---------R = 1.6 x 106 to 4.4 / i 0 I r 0L,l :

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Figure 2.- Concluded.

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Unflagged symbols denote smooth condition 4;.

Flagged symbols denote leading-edge roughness l &m R = 0.8 x 106 to 1.e x 106 ‘ i I r7 9 b 1 0 1.0 1.15 ------_ _ R = 1.5 x lo6 to 4.0 x lo6 1;.

‘ i.0 * I ‘ $ {: cl ,j ‘ : [ 1.2 1.0 -1.2 a 1.0 !I ’ ) b rl.2 _ .8 0” ‘ i J I c, 1.0 ..-A .rl .8 i?

.8 1.2 .6 b 1.0 .8 - .6 Mach number, M Mach number, M (a) Section lift coefficient.

Figure 3.- Aerodynamic characteristics at various angles of attack obtained

with l.O-foot chord NACA 64,-013 airfoil section.

- Unflagged symbols denote smooth condition Flagged symbol .s denote leading-edge roughness .

I’ .

R = 0.8 x 106 to 1.8 x 106 ;:..

R = 1.5 x 106 to 4.0 x 106 i.

B ‘ 0 Sk . . . .8 b.m* 1;: .7 .6 -.

.6 .7 - - .6 - .5 - .6 - - .08 oa I .07 - m a ..-I .06 .d f ‘ ) - fz d .05 ” - al .d 0 .04 -2 .04 Fi D Ei - .03 .rl * .- -s SO3 - .03 - V ,02 - c ia .02 r .Ol - .02 = .- - 1.02 1.Ql q z .Ol

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+ - .6 .7 .8 .9 . 5 .6 .7 .0 ’ ’ .9 Mach number, M Mach number, M (b) Section drag coefficient.

Figure 3.- Continued.

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,Unfla&wJ symbols denote smooth condition Flagged symbols denote leading-adie rough ess R = 0.a x 1 0 6 to 1.8 x 1 0 = 1.5 x 106 to 4.0 x 106 ----------R

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.4 .5 .6 .-I A .2 .3 Mach number, Y Mach mmber, Y (c) Section quarter-chord.pitching-moment coefficient.

FQure 3.- Concluded.

I-

(Jnflagged symbols denote smooth condition = Flamed symbols denote leading-edge roughness .

= R = 0.8 x lo6 to 2.6 x lo6 . .

_----__--_ R = 1.6 x IO6 to 4.0 x JO6 . .

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5’ .6 .2 a

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\, -- .8 v .6 -.

.6 .7 ,0 .9 .2 .3 .4 .5 .2 .3 .4 .5 .6 i7 .a .9 Mach number, M Mach number, M (a) Section lift coefficient.

Figure 4.- Aerodynamic characteristics at various angles of attack obtained

with l.O-foot chord NACA 643-018 airfoil section.

Unflagged symbols denote ! ?a 100th condition Flagged symbols denote 1.ead .ing-edge roughness R = 0.8 x 106 to 2.6 x 106

-----R-

------

= 1.6 x 106 to 4.0 x 106 .0; 7 .0t de .Ot: i ’ L .OL .02 I

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, .0: .Ol l- .02 q ( 1 I .Ol > .OS L 00 .O! 1 ( 1 L- .2 .3 .4 .5 .6 .7 .8 .9 .2 .3 .4 .5 .6 .7 .8 .9 Mach number, M Mach number, M (b) Section drag coefficient.

Figure 4.- Continued.

0 I . a . .

Unflagged symbols denote smooth condition Flagged symbols denote leading-edge roughness .l R = 0.8 x 1~6 to 2.6 x 106 ------- R = 1.6 x 106 to 4.0 x 106 .l

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Figure 4.- Concluded.