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4

0. OATIONAL ADVISORY COMMITTEEFOR AERONAUTICS

C TECHNICAL NOTE 3172

EFFECTS OF LEADING-EDGE RADIUS AND MAXIMUM THICKNESS-

CHORD RATIO ON THE VARIATION WITH MACH NUMBER

OF THE AERODYNAMIC CHARACTERISTICS OF

SEVERAL THIN NACA AIRFOIL SECTIONS

By Robert E. Berggren and Donald T. Graham

Ames Aeronautical LaboratoryMoffett Field, Calif.

Washington

April 14, 1954

4 'i

IF NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

TECHNICAL NOTE 3172

EFFECTS OF LEADING-EDGE RADIUS AND MAXIMUM THICKNESS-

CHORD RATIO ON THE VARIATION WITH MACH NUMBER

OF THE AERODYNAMIC CHARACTERISTICS OF

SEVERAL THIN NACA AIRFOIL SECTIONS1

By Robert E. Berggren and Donald J. Graham

SUMMARY

A wind-tunnel investigation has been made to .eterming the effectsof leading-edge radius and maximum thickness-chord ratio on the variationwith Mach number of the aerodynamic characteristics of several thinsymmetrical NACA 4-digit-series airfoil sections. The Mach number rangeof the investigation was from 0.3 to approximately 0.9 and the corre-sponding Reynolds number range from approximately 1 X 106 to 2 x 106.

The variations with Mach number of the lift, drag, and pitchingmoment for a 4-percent-chord-thick airfoil section are not significantlyaffected by a change of leading-edge radius from 0.18 to 0.53 percent ofthe chord. A similar conclusion can be drawn for a leading-edge-radiusvariation from 0.10- to 0.40-percent chord on a 6-percent-chord-thicksection.

Progressive improvement of the variation of lift-curve slope withMach number, the lift and drag-divergence characteristics, and the max-imum section lift characteristics at Mach numbers above 0.6 results fromreduction of the maximum thickness-chord ratio from 10 to 4 percent.Section pitching-moment characteristics are not greatly affected by vari-ation of the maximum thickness-chord ratio.

INTRODUCTION

To investigate the influence of airfoil leading-edge radius on thevariation with Mach number of the aerodynamic characteristics of thin

1 Supersedes NACA RM A50D04, "Effects of Leading-Edge Radius and MaximumThickness-Chord Ratio on the Variation With Mach Number of the Aero-

A dynamic Characteristics of Several Thin NACA Airfoil Sections" byRobert E. Berggren and Donald J. Graham, 1950.

~/

2 NACA TN 3172

airfoil sections, a series of airfoil tests was conducted in the Ames 1-by 3-1/2-foot high-speed wind tunnel. The results of the investigationfor a thickness-chord ratio of 10 percent have been reported in refer-ence 1. The results for thickness-chord ratios of 6 and 4 percent arereported in the present paper. The basic thickness form of the airfoilsinvestigated was the NACA 4-digit series (see reference 2) with meximumthickness at 40 percent of the airfoil chord.

In addition to the leading-edge-radius study, the investigation per-mitted further analysis of the effects of thickness-chord-ratio variationon the characteristics of airfoil sections at high subsonic Mach numbers.This analysis is also contained in the present report.

NOTATION

a° section lift-curve slope, per degree

c airfoil chord, feet

cd section drag coefficient

cl section lift coefficient

Imax maximum section lift coefficient

c '/ section pitching-moment coefficient about the quarter-chord point

M free-stream Mach number

M Mach number for drag divergence, defined as the Mach number at

(dcd\whichk-c = 0.1

\dM/

CI= constant

M1 Mach number for lift divergence, defined as the Mach number at

which( d c =0\dM o -constant

V free-stream velocity, feet per second

4- -~---- *~

NACA TN 3172 3

v local velocity, feet per second

Ava increment in local velocity corresponding to additional type of

load distribution, feet per second

x distance along chord from leading edge, fraction of chord

y distance perpendicular to chord, fraction of chord

CL section angle of attack, degrees

DESCRIPTION OF AIRFOILS

The airfoil sections of the present study are:

leading-edge radius

NACA airfoil designation (percent chord)

0004 - 1.10 40/1.575 0.180004 - 3.30 40/1.575 .53

0006 - 1.10 40/1.575 .40

0006 - .70 40/1.575 .25

0006 - .27 40/1.575 .100008 - 1.10 40/1.575 .700010 - 1.10 40/1.575 1.10

The first digit of the airfoil designation indicates the camber in per-

cent of the chord; the second, the position of the camber in tenths of

the chord from the leading edge; and the third and fourth, the maximum

thickness in percent of the chord. The decimal number following the dash

is the loading-edge-radius index; the leading-edge radius as a fraction

of the airfoil chord is given by the product of the radius index and the

square of the thickness-chord ratio. A radius index of 1.10 is standard

for the NACA 4-digit-series airfoil sections. The two digits immediately

preceding the slant represent the position of maximum thickness in per-

cent of the chord from the leading edge. The last decimal number is the

trailing-edge-angle index, the angle being twice the arc tangent of the

product of the angle index and the thickness-chord ratio.

The coordinates of the airfoils investigated are given in tables I

to VII. The profiles are illustrated in figure l and the theoretical

low-speed pressure distributions, determined by the method of reference 3, ! -

in figure 2.

- -0

4-

4 NACA TN 3172

APPARATUS AND TESTS

The tests were made in the Ames 1- by 3-1/2-foot high-speed windtunnel, a low-turbulence two-dimensional-flow wind tunnel.

The airfoil models, constructed of aluminum alloy, were of 6-inchchord and completely spanned the 1-foot dimension of the wind-tunnel testsection. End leakage was prevented by means of contoured sponge-rubbergaskets compressed between the model ends and the tunnel walls.

Measurements of lift, drag, and pitching moment were made at Machnumbers from 0.3 to as high as 0.9 for each of the airfoils at angles ofattack increasing by 10 or 20 increments from .2o to a maximum of 12° .This range of angles of attack was sufficient to encompass the lift stallup to Mach numbers of the order of 0.8. The Reynolds number of the testsranged from approximately 1 x 10e at the minimum Mach number to approx-imately 2 x 10e at the highest Mach numbers.

Lift and pitching moments were evaluated by a method similar to thatdescribed in reference 4 from integrations of the pressure reactions onthe tunnel walls produced by the airfoil models. Drag measurements weremade by means of wake surveys using a rake of total-head tubes.

RESULTS AND DISCUSSION

Section lift, drag, and quarter-chord pitching-moment coefficientsfor the airfoil sections investigated are presented as functions of Machnumber at constant angles of attack in figures 3, 4, and 5, respectively.The characteristics for the 10-percent-thickness-chord ratio are takenfrom reference 2. The angles of attack indicated in the figures repre-sent but nominal values, being subject to a maximum experimental errorin setting of 0.150. The characteristics have been corrected for tunnel-wall interference by the methods of reference 5. Dashed lines have beenused in the figures to indicate the region of possible influence of wind-tunnel choking effects on the results.

leading-Edge Radius

Within the limits of the present investigation, the leading-edgeradius does not significantly influence the variation with Mach numberof the aerodynamic characteristics of 4- and 6-percent-chord-thick air-foil sections. A small superiority in the maximum section lift coeffi-cient at Mach numbers from 0.4 to 0.75 is indicated in figures 6 and 7for the 4-percent-thick airfoil wJth the very large nose radius. For the

-,

NACA TN 3172 5

6-percent-thick sections no important differences exist. No importanteffect of nose radius change on the lift-curve-slope variation with Machnumber is indicated in figure 8 for either thickness-chord ratio. Theminimum drag coefficient is noted from a study of figure 9 (illustratingthe variation of section drag coefficient with section lift coefficientat constant Mach number) to be lower at all Mach numbers for the 4-percent-thick section with the standard leading-edge radius; but, at moderate tolarge lift coefficients for Mach numbers up to 0.7, the drag coefficientsare lover for the profile with the larger nose radius. The latter trendcan also be noted from this figure for the 6-percent-thickness-chord ratio.No real differences are observed in the variations of section pitching-moment coefficient with section lift coefficient at constant Mach number(fig. 10) for the sections with the various leading-edge radii.

Maximum Thickness-Chord Ratio

A progressive improvement in airfoil-section lift characteristicsresults from reduction of the airfoil maximum thickness-chord ratio from10 to 4 percent. From figure 11, the lift-divergence Mach number isobserved to increase nearly linearly with thickness reduction.2 Figure12 illustrates the gain in maximum section lift coefficient with decreasein maximum thickness at Mach numbers above 0.6. The values at Mach num-bers below about 0.6 are subject to question because of the low scale.However, the results of the investigation of reference 6 indicate that atthe higher Mach numbers the values are not much influenced by the rela-tively low test Reynolds numbers (approximately 2X106). The effects ofmaximum thickness-chord-ratio variation on the section lift-curve slope,illustrated in figure 13, are what should be expected in that each succes-sive reduction of thickness increases the Mach number at which the lift-curve slope breaks.

The effect of reduction of thickness-chord ratio on the Mach numberfor drag divergence (fig. 14) is to increase markedly the value of thisparameter at zero lift. With increasing lift coefficient this favorableeffect diminishes, becoming very small at a lift coefficient of 0.5.

At Mach numbers below 0.7, the variation of section drag coefficientwith section lift coefficient (fig. 9) is adversely affected by reductionof the maximum thickness; for Mach numbers greater than 0.75, the converseis true. The minimum drag coefficient is progressively decreased withmaximum thickness reduction at all Mach numbers.

Maximum thickness-chord ratio, within the limits of the presentinvestigation, has no important influence on airfoil-section pitching-moment characteristics.2The portions of the curves shown for the lift-coefficient range fromapproximately -0.2 to 0.2 represent estimated values of the lift- 4

divergence Mach number, there being insufficient data to permit positivedetermination of this parameter near zero lift. The lift-divergenceMach number, of course, has no significance at zero lift.

S IVA

6 NACA TN 3172

CO CLUSI0NS

From the results of a high-speed wind-tunnel investigation of the

effects of leading-edge radius and maximum thickness-chord ratio on thevariation with Mach number of the aerodynamic characteristics of severalthin symmetrical NACA 4-digit-series airfoil sections, it is concluded:

1. The variations with Mach number of the lift, drag, and pitchingmoment for a 4-percent-chord-thick airfoil section are not significantlyaffected by a change of the leading-edge radius from 0.18 to 0.53 percentof the chord. The same is true for a leading-edge-radius variation from0.10- to 0.40-percent chord on a 6-percent-chord-thick section.

2. Reduction of the maximum thickness-chord ratio from 10 to 4percent progressively improves the variation of lift-curve slope withMach number, the lift and drag-divergence characteristics, and the max-imum section lift characteristics at Mach numbers above 0.6.

3. Section pitching-moment characteristics are not greatly affectedby variation of the maximum thickness-chord ratio.

Ames Aeronautical Laboratory,National Advisory Committee for Aeronautics,

Moffett Field, Calif., May 4, 1950

REFERENCES

1. Summers, James L., and Graham, Donald J.: Effects of Systematic

Changes of Trailing-Edge Angle and Leading-Edge Radius on theVariation with Mach Number of the Aerodynamic Characteristicsof a 10-Percent-Chord-Thick NACA Airfoil Section. NACA RMA9G18, 1949.

2. Stack, John, and von Doenhoff, Albert E.: Tests of 16 RelatedAirfoils at High Speeds. NACA Rep. 492, 1934.

3. Theodorsen, Theodore: Theory of Wing Sections of Arbitrary Shape.NACA Rep. 411, 1931.

4. Abbott, Ira H., von Doenhoff, Albert E., and Stivers, Louis S., Jr.:Summary of Airfoil Data. NACA Rep. 824, 1945. 1

NACA TN 3172 7

5. Allen, H. Julian, and Vincenti, Walter G.: Wall Interference in aTwo-Dimensional-Flow Wind Tunnel, With Consideration of the Effectof Compressibility. NACA Rep. 782, 1944.

6. Spreiter, John R., and Steffen, Paul J.: Effect of Mach and ReynoldsNumbers on Maximum Lift Coefficient. NACA TN 1044, 1946.

4'

.. . . . . . . . . . .

8 NACA TN 3172

TABLE I. - COORDINATES AMD THEORETICAL PRESSURE DISTRIBUTIONS FORTHE NACA 000o4-i.o 40/1.575 AIRFOIL

Y 2(percent c) (percent c) (v/V) v/V AVa/V

0 0 0 0 5.5651.25 .605 1.125 I.061 1.4272i5 .818 1.130 1.063 1.0105.0 1.090 1.123 1.06o .7057.5 1.270 1.115 1.056 .56610 1.413 1.1o8 1.053 .48315 1.620 1.099 I.o48 .38220 1.765 1.097 1.048 .32030 1.940 1.093 1.046 .24340 2.000 1.088 1.o43 .19550 1.940 i.085 1.042 .15660 1.773 1.082 i.040 .12570 1.493 1.061 1.030 .09680 1.106 1.032 i.016 .06990 .622 .994 .997 .03795 •342 .940 .970 .013

100 .040 0 0 0

L. E. radius: 0.18 percent c.

- N

2F

NACA TN 3172 9

TABLE II. - COORDINATES AND THEORETICAL PRESSURE DISTRIBUTIONS FORTHE NACA 0004-3.30 40/1.575 AIRFOIL

x y 2 V/V(percent c) (percent c) (v/V) Ava/V

0 0 0 0 3.5151.25 .931 1.328 1.153 1.1992.5 1.196 1.317 1.148 .9945.0 1.468 1.214 1.102 .6857.5 1.611 1.182 1.087 .549

10 1.717 1.153 1.074 .46615 1.799 1.112 1.054 .36620 1.862 1.091 1.045 .30630 1.955 1.078 1.038 .23440 2.000 1.082 1.040 .18950 1.940 1.083 1.041 .15460 1.773 1.079 1.039 .12670 1.493 1.059 1.029 .09980 1.106 1.033 1.016 .07590 .622 .994 •997 .04995 .342 .935 .967 .033

100 .040 0 0 0

L. E. radius: 0.53 percent c.

IJ

10 NACA TN 3172

TABLE III. - COORDINATES AID THEORETICAL PRESSURE DISTRIBUTIONS FORTHE NACA 0006-1.10 40/1.575 AIRFOIL

x y v/V

(percent c) (percent o) (/)vV

0 0 0 0 3.7811.25 .907 1.149 1.072 1.3612.5 1.228 1.174 1.084 .9745.0 1.633 1.174 1.084 .6847.5 1.908 1.164 i.080 .551

10 2.120 1.158 1.076 .47015 2.433 1.145 1.070 .37220 2.645 1.141 i.068 .31230 2.915 1.143 1.069 .23940 3.000 1.141 1.068 .18950 2.915 1.128 1.062 .15460 2.660 1.115 1.056 .12570 2.240 1.088 1.043 .09880 1.66o 1.053 1.o26 .0729o •934 1.002 1.001 .04595 .514 .915 •957 .027

100 .060 0 0 0

L. E. radius: 0.40 percent c.

NACA TN 3172 Ij

TABLE IV. -COORDINATES AND THEORETICAL PRESSURE DISTRIBUTIONS FOPTHE NACA 0006-0.70 40/1.575 AIRFOIL

x y (Vlv) VlV iv(percent c) (percent c)

0 0 0 0 4.5201.25 .766 1.083 1.041 1.3652.5 1.067 1.123 1.060 .9775.0 1.473 1.138 1.067 .6kS77.5 1.767 1.142 1.069 .555

10 1.989 1.144 1.069 .47415 2.354 1.148 1.072 •37720 2.607 1.152 1.073 .31730 2.908 1.148 1.072 .24140 3.000 1.145 1.070 .19350 2.915 1.136 1.066 .15660 2.660 1.117 1.057 .12670 2.240 1.095 1.046 .0)980 1.660 1.056 1.028 .07490 .934 .997 .999 .04795 .514 .924 .961 .030

100 .060 0 0 0

L. E. radius: 0.25 percent c.

II

12 NACA TN 3172

TABLE V. - COORDINATES AND THEORETICAL PRESSURE DISTRIBUTIONS FORTHE NACA 0006-0.27 40/1.575 AIRFOIL

x y(percent c) (percent c) (v/V) 2 v/V AVa/V

0 0 0 0 6.8931.25 .566 .962 .981 1.3492.5 .840 1.029 1.015 .9745.0 1.247 1.077 1.038 .6897.5 1.567 1.097 1.047 .558

10 1.826 1.119 1.058 .48015 2.246 1.140 i.068 .38320 2.546 1.154 1.074 .32230 2.900 1.160 1.077 .24540 3.000 1.148 1.071 .19450 2.914 1.134 i.065 .15760 2.660 1.120 1.058 .12770 2.240 1.097 1.047 .09980 1.660 1.058 1.029 •07390 •934 .999 .999 .04695 .514 .920 .959 .028

100 .060 0 0 0

L. E. radius: 0.10 percent c.

"I

NACA TN 3172 13

TABLE VI. - COORDINATES AND THEORETICAL PRESSURE DISTRIBUTIONS FORTHE NACA 0008-1.10 40/1.575 AIRFOIL

x y

(percent c) (percent c) (v) 2 v/V

0 0 0 0 2.9231.25 1.210 1.138 1.067 1.3292.5 1.636 1.228 1.108 .9745.0 2.179 1.236 1.112 .6867.5 2.540 1.223 1.106 .552

10 2.825 1.217 1.103 .47115 3.240 1.206 1.098 .37420 3.530 1.199 1.095 .31430 3.889 1.194 1.093 .23940 4.000 1.191 1.092 .19150 3.889 1.182 1.087 .15560 3.545 1.160 1.077 .12570 2.985 1.123 1.060 .09880 2.212 1.075 1.037 .07290 1.243 .994 .997 .04595 .684 .919 .958 .029100 .080 0 0 0

L. E. ridius: 0.70 percent c.

4

I

14 NACA TN 3172

TABLE VII. - COORDINATES AND THEORETICAL PRESSURE DISTRIBUTIONS FORTBE NACA 0010-1.10 40/1.575 AIRFOIL

x y2(percent c) (percent c) (v/V) v/V AVa/V

0 0 0 0 2.3241.25 1.511 1.1o8 1.053 1.2862.5 2.044 1.245 1,116 .9665.0 2.722 1.286 1.134 .6907.5 3.178 1.277 1.130 .556

10 3.533 1.269 1.127 .47515 4.056 1.261 1.123 .37720 4.411 1.248 1.117 .31630 4.856 1.244 1.116 .24140 5.000 1.242 1.115 .19350 4.856 1.231 1.110 .15560 4.433 1.211 1.101 .12670 3.733 1.155 1.074 .09880 2.767 1.089 1.043 .07290 1.556 .980 .990 .04595 .856 .912 .955 .030100 .100 0 0 0

L. E. radius: 1.10 percent c.

ku --.

PTACA TN 317P 15

NACA 0004-1.10 401.575

NACA 0004-5.30 401.575

NAGA 0006 -1l0 4 011.5 75

NACA 0006-0.70 4 0/1.5 75

NAcA 0006-0.2 4011.5 75

NAGA 0008 -. 10 4011.575

NACA 00/0-1/10 401.5 75

Figure Z.- NACA air foil profiles Investigated.

16 NACA TN 3172

-.

e

-

z

'

"\

I Ib

1I

I.

'.44

i3FF ACA TN 3172 17

I--

I 0

cm

q)

18 NACA TN 3172

- -- -- - U -O -

gdizz %

MACA MN 3172 19

1.6

~ -2*-/. A

0 06*

44

I ~--- -------- - --C

.2S

.3 . .5 6 . .8 9 1

20 NACA TN 3172

1.2-

40'

-6 --

til

'10

-6-

*.:2 .3 .4 .5 .6 .7 .8 .9 /.0Mach rumber, M

(b) NACA 0004-.30 40/1575 Airfoil.

Figure 3.- Conived.

NACA TN 3172 21

1.2-

.-

.6 6-2

V62*

'/0

.2 . .4 .5 .6 .7 .8 .9 /.0Mach nme, M

(c) NACA 0006-110 40/1575 Arfoil.

Figure 3.- Continued.

22 NACA SW 3172

1.2

.8- -- - -- - -

.6

.2 84 .5. 7.8.

Mach numberv M

(d) NACA 0006 -0.70 401.575 Airfo/.

Figmr 3.- CmAmokwd.

NACA TN 317r2 23

/.0-----

off

-6~2 6 ~ --

c 2" ----

M4hnmbr

(e6AA00-.7 017:i~i

Pwe3- onf8 ud

v 120 ........

r~e

.2-

24 NACA TN 3172

1.2

CCI

.3 .4 . 6 .7 . 9 1

Mac nmbe,00 0

(f NAA00-11 0175Arol

Fiue3-Cc4/ud

0 2*

Z 40

v 60

IFTACA TIT 3172 25

.4

V 6

120L1

.2 .5 .4 .5 .6 .7 .8 .9 1.0Mach number M

(g) NACA 00/0-/*/0 40/1575 Airfoil,

Figure 3.- Concluded.4

26 NACA TN 3172

.20-

de6-- - - -- -

.14 -

0012

. 0 6 - ------,:-

.06

(a) NACA 0004-10 40/1.57'5 AirfoilFiue4.- Variation of secton drag coefficient with Mach number at .

constant angles of attack. ,"

/0 /a

I\

NACA TN 3172 27

.--

.16--- - -

L~ 40

. 0 / <- -/0

'106,

.064-- - - - -- -

01

Mach niumber, M(b) NVACA 0004-3.30 40/1575 Airfoil.

Figure 4.- Continued

4'

28 rUAVA TN 37

.22

16*

.14 K-2

' 1 002 7- - - - -----

4-.4/0$ 0

- Y.~-

0r

Cz,

NACA TN 31(2 29

.2 21---

I'd

.12

o 00 -

/0 o /*

A 4

.08 -> 80--1/0

.06---

.04------

.02

02 .3 .4 .5 .6 .7 .8 .9 10

Much mumber, M(d) NAGA 0006-0.70 40/1575 Airfoil.

Figure 4.- Coninued.

! IVY

30 NACA TN 3172

.20-

.18------------- -

.12 - - -

0 00

L4*

.06

.04_

02 .3 .4 .5 .6 .7 .8 .9 i0Mach numberm A

(e) NA CA 0006-0.27 4011.575 Airfoil.Figure 4. - Continued.

NACA TN 3172 31

.22

.20-------------------

.16 -- - --- -

14

0

V 61 80

06 o -0

(V4-

0-5 .6 . .8 .9 z 0Mach number, M

(fJ NASA 0008-110 4011575 Airfaf/.Figure 4.- Continued.

32 NACA TN 31.72

.22---

.20- - - --- --

o 0.

A4

'/0*

-i2-

.5 . .8 .9/

NACA TN 3172 33

.4

4 oO

76/0 *--------- --

.2 . .4 .5 .6 .7 .8 .9 /.0Mach number, M

()NACA 0004-110 401.575 Airfoil,

Figure 5.- Variation of sectlon pitching-moment coefficiet wit/i Machnumber at constant angles of attack.

34 NACA TN 3172

.4.

-3 -2

000

S.20 2: -

z.4*

/,012IR

-0 --

-4 .--- 4-- 5-----7--8--9---

Mach number, M

(b) A'ACA 0004-5.30 40/1.575 Airfoi.

Figure 5.- Continued.j

NACA TN 3172 35

0 0

.2 2"

V 6"

(00

,04

.. .. I '

Mach number, M

(C) NACA 0006-1IO 40//.575 Airfoi A

t Figure 5.- Continued.

A .

r

36 NACA TN 3172

.4 E

0 00

*~.2 24L4*v60

A!A

.2V

NACA TN 3172 37

0 0O, O"

ve .2~ 0 -------

D0

". 3 4 5 6 .7 .8 .9 1.0

Mach number, M

4(6) NACA 0006-0.27 40/1.575 Airfoil

Figure 5.- Con/inued.,

414

.2 .3 .4 . .6 7 .8 .9 1

38 NACA TN 3172

.4

0 00

.2 2~ A40

.b 76

2 - --

* 2 .3 .4 5 .6 7 .8 .9 1

Mach number, M

(f) fVACA 0008-110 40/1575 Airfoil,

Figure 5.- Continued.

NACA TN 3172 39

! ~.4 |...

0o 0"0 / 0

S.2 0 --=A 40

v6

v 120

- _ -

bb

~ 0

.4 - --- - - - - - - - - - - - -

.2 .3 .4 .5 .6 .7 .8 .9 /0Mach number, M

(g) NACA 0010-IO 40/1.575 Airfoil.

Figure 5.- Concluded

,b .- - - --- - - -'4

40 NACA TN 3172

QN

IkI

C'I-

INS%

9iey~o ji W 'o m9

4 . i* I

3F

DIACA TNf 3172 4

HT,-

...... . ....

ra -'-

42 NACA TN 3172

FV

NACA TN 3172 43

NJN.

n I iz

FI

lb0

to%

%J %~ ~ ,. IWai2 wo all.3JJO 411/ 109

44 NACA TN 3172

--

E

1.7 W vwoo mll lwoo

ITACA TN 3172 45

III

Cl4

.4-

VV

46 NACA TN 3172

-J41

I T

Q03 sD3JJO 41 O/G

Wavvip- all q/10S -

NACA TN 3172 47

1.2NAGA 0004-3.30 40/11523

NACA 0004-110 40/11575

10

Mach number, M.6 .

(G) Maximum thickness - chord ratio of 0.04.

1.0--------- -- -

.8

b" 6~ ____NAGA 0006 - 1/0 401.575 .

-__NAGA 0006 - 070 40/1575NACA 0006-0.27 40/1575

.3 .4 .5 .6 .7 .8 .9Mach nme, M

(b) Maximum thickness - chord ratio of 0.06.

Figure 7- Effect of change of leading-edge radius on the variationof maximum section lift coefficient with Mach number.

"'dWYK

48 NACA TN 3172

.4NACA 0004-3.0 4011575

NACA 0004-1.10 40/1575

0

.2 .3 .4 .5 .6 .7 .8 .9

Mach number, M

(a) Maximum thickness -chord ralio of 0.04.~.4

___NACA 0006-ZI0 40/1575NAGA 0006-07040/1575

-- NAGA 0006-02740/1575I .3

.2 ---------------

C - - - - -I I I

.2 .3 .4 .5 .6 .7 .8 .9Mach number, M

b) Maximum Ihickness - chord rato of 0.06.

Figure 8- Effed of change of leading-edge radus on the variaion ofsecIon lift-curve slope with Moch nWber.

! t

7FNACA TN 3172 49

INI'

.875

.85 16 -d1 2 . 6 . .

.75.

42 VIA1

50 NACA TN 3172

e757

.e5 -- T 0"A

.1 s.825 A-

.7.4 50- .751 YAl I I

(b *75 0- 3.0 4/.7 AlIolFgr 9 . 5 Co 1 .i/ed.il

70-

16 .70

67-

NACA TN 3172 5

I I3

'I. __

'.875

.85

'3 __0875

'3 ~ z~z~m85

.8 .04 .5.77

-8-~ ~4- 02 7.5.0£

.14 NAA0066704/175irol

'11.5

-

52 NACA TIT 3172

.875 1 \ l

.85-

.85

.4.75

.070

.16 44.L0Q .142.64 0.r. .6.5 WV/

IS i. 9-Cntne

NACA TN 3172 53

.875-

N I

I'I

AS --

.060

.04 .50- = -

.775 .750-21

.2046. 70 74W

It AM

NACA TN 31.72

'.I5

4%44

.08 .65

'a.06 .5

* ,04 .50

02 .40

JO8 6 -.4 -.2 0 .2 W4. 8 £ 1£2Secam li~ ft oeific/ealf, c1(f) NACA 0008-110 401575 A/dfoll.

Figure q.- Conalned

NACA TN 3172 55

1 .87 - 1 1

1-7 -t

.4975 - .82 - -- - --

.85-

t.77

J. I7I2I4 8 .754~20246.10£

J67

.12 -

.65 $25

56 NACA TN 3172

87 75

.855

.85 I a -

.805 -

.77 5- ---- - -

'.7075 L It.65 -- I J- ---- -' - -

.1.75--•.5 ---- ' - a I/

.60

~~ ~1.675-- I :::.65 "

.625 ..

.600-1

lot a .5"---: ow m"\E

.2 .50 . - "

I .40 "-40 - - --

iMir € J..: AAt 0 ~~~30 30 -- a a

-.8 -.6 -.4 -2 0 .2 .4 .6 .8 1.0 Z2

Section lift coefficint, 1c

(a) NACA 0004-1.10 40/W575 Airfoil.

Figure /0.- Var/ation of section pitchIng-moment coefficient withsection ift coefficient of constant MOh rwmbers.

3FITACA TN 3172

5

~ 1 .0- 0-I - - +

.80

.77

NACA TN 3172

.85

I.85

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~ .LOW..67

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NACA TN 3172 59

.875-

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Fi~v /. 75 ontiwe

60-~

60 NACA TN 3172

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S.1 .4o T:L'IJ.4

so -3 M13-t- m

Section lift coeffient, czjS

(e) NACA 0006-0.27 40/1.575 Airfoil.

Figure /0.- Continued.

J£

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NACA TN 3172 61

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(f) NACA 0008-I/O/ 40/1,575 Airfoil.

d .40

Figure /0. - Continued.0 .5-

u .

4.'

62 NACA TI 3172

-- .87*- a-aaa--: :• -7- - -s-

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q~~t ~ 55 -

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Section lift coefficlent, c/

(g) NAGA 0010-1.10 40/A.575 airfoil.

FIgure /0.- Concluded. 4

ITACA TIN 3172 63

V It

// I / I I I :.II

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, / I"

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64 NACA TN 3172

/.2_

.0

.6 ____NACA 0008-1.10 4011.575MACA 0006-110 4011.575NACA 0004-1.10 40/1575

.3 .4 .5 .6 .7 .8 .9Mach number, M

Figure /2.- Effect of change of maximum thickness- chord ratio onthe variation of maximum section lift coefficient With Mach number

__4_ NAGA 00/0-/0 40/1575NACA 0006-1/0 4011.575- - - -

NACA 0004-110 40/1575

- - NG 004-1.10- ---- ---- --- 4~O

.2 .3 .4 .5 .6 .7 .68.Mach number, M

Figure /3. - Effect of change of maximum thikness - chord ratio on thievariation of section lift-curve slope With Mach number

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fb

th

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