L-

COPY r-2 r

RM A9G19

RESEARCH MEMORAND

. COMPARISON OF THE AERODYNAMIC CHARACTERISTICS OF

3 THE NACA 0010 AND 0010-64 AIRFOIL SECTIONS b

s CJ AT HIGH- SUi3SONiC M&H NUtiBFsRS

By Perry P. Polentz

Ames Aeronautical Laboratory Moffett Field, Calif.

-P 8 NATIONAL ADVISORY COMMITTEE +%I -(I;

3 FOR AEROtiAUTlCS M Q*

TX WASHINGTON

$2 October 7,1949

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NAC!A RM AgGIg

NA!JlIONLLLADVISORYCMMITTE33FOR AEWRiAUTICS

RESEiARCHMEMOJ3ANDUM

CCMPARISON OFTHEAERODYNAMIC CHARACTERISTICS

THENACA OOlOAND OOlti4AIRFOILSEZ'TIONS

AT HIGHSUBSONICMACHmERS

By Perry P. Pole&z

SUMMARY

OF

A wind-tunnel investigation has been conducted to determine the lift, drag, and pitchi- nt characteristics of the NAC!A 0010 and 001&64 airfoil sections at Mach nu&ers up to 0.91 and Reynolds nuuibers betweenl.0 x106 and 1.9 X 1Oe. The results are compared to Illustrate the effects of varying the chordwise location of maximum thiclmess from 30-percent to k&percent chord on the principal high- speed characteristics of the sections.

A virtually unchanged Mach number for lift divergence, a decrease in ILf-kxrve slope of approximately 10 percent, and a reduced maximum lift coefficient at Mach mmibers below 0.70 were associated with the more rearwar d location of maximum thickness. The Mach number for drag divergence was fncreased about 0.05 at lift coefficients up to 0.4, but the rate of drag rise above the Mach number for drag divergence was notappreciablychanged. Pitching moment was affected to a negli- gible degree.

INTRODUCTION

The chszacteristics at high Mach numbers are available for relatively few airfoil sections of the NACA &digit series. The present experimental investigation.was undertaken to obtain such data for the NACA 0010 and 001064 airfoil sections at Mach nmibers rang- up to 0.91. A further purpose was to appraise the effect of varying the position of mximum thictiss frcan jC+percent chord for the NACA 0010 profile to hercent chord for the NACA 0011>-64 profile.

NOTAJXON

a0 section Iff-ticurve slope, per degree

Cd section drag coefficient

NACA RM AgGlg .

section lift coefficient

maximum section lift coefficient

section quarteMhord pitching-moment coefficient

Mach number of free stream

Mach number for drag divergence

Mach number for lift divergence

section angle of attack, degrees

APPARATUS AND TESTS

The tests were conducted in the Ames l- by j-l/&foot high-speed wind tunnel. Thistunnelisatwo-dime nsional, close&return type having a rectangular test section of the indicated cross-sectional dimensions, and is provided with sufficient power to obtain choked flow in the presence of any model. Its contraction ratio is 16,l:l. Atmospheric air, maintained at barometric pressure in the settling chamber, forms the working substance.

Lift and pitc went data are obtained by use of a method similar to that described in reference 1 from measurements of the reactions on the tunnel floor and ceiling of forces experienced by the airfoil. Drag is determined from w&s-survey measurements made with a rake of total--head tubea. By use of these methods it is possible to seal completely the gap between the sides of the tunnel and the ends of the airfoil, and ensure that two4imensional flow is obtained over the entire surface without interference with force measurements.

Scale drawings of the profiles tested are reproduced in figure 1, and the corresponding coordinates are tabulated in table I, from which it will be noted that the point of maximum thiclmess is located at 3+percent chord for the NACA 0010 airfoilsection and at 4.0 percent for the NACA 001&64. (The significance of the airfoil notation used is explained in reference 2.) The chord length employed for-the tests x&6 6 inches; the models were mounted at the center line of the tunnel and spanned the L-foot dimension. The airfoils were fabricated of aluminum alloy, the deviation from nominal dimensions being held to 0.00~inch maximum. All surfaces were carefully polished to a mirrorlike finish.

The Mach number of the tests was varied from 0.3 minimum to a maximum value lying between 0.75 and 0.91, the exact range depending upon the angle of attack but being sufficient to encompass the lift stall up to Mach numbers of the order of 0.8. Data were secured for angles of attack between -2O and l2O at increments of 2', and at -lo

4

NACA RM Awl9 3

.

and 10. Reynolds numbers of the investigation varied from 1.0 X log nl3dmlm to 1.9 x lo= maxinlum.

The results obtained in the wind tunnel have been corrected by the method of reference 3 to account for the constriction fn the chasnel caused by the model andbythe wake. The magnitude of these corrections increases both with Mach number and with szgle of attack, but, ti general, smounts to less than 2 percent of the values reported. This same reference demonstrates that no correction is possible for data obtatied at the choking Mach number. Dashed lines are used on the figures to indicate measurements made in the vicinity of this Mach nuxiber which are of doubtf'ul validity.

Lift, drag, and pitching+Ement coefficients for the NACA 0010 and 001044 airfoil sections are presented as functions of Mach nmiber in figures 2, 3, and 4. An indication of the accuracy of the lift and pitch-n-t measurements is afforded by the symmetry of the curves at low lift coefficients. Owing to such variables as stream angularity, model asymmetry, and errors in setting the sngle of attack in the xind tunnel, discrepancies equivalent to as much as 0.2' in angle of attack nay be observed.

Lift as a function of Etngle of attack is shown for various Mach numbers for the two afrfoil sections in figure 5. Comparison of the priz+ cipal lift parameters is provided in figures 6 and 7. Figure 6 discloses that no significant difference exists for the lift-divergence Mach nu&ers of the two profiles. (Lif-t+divergenceMachnuuiberis arbitrarily defined as that Mach number at which the first point of inflection occurs in the lift coefficient versus Mach nusiber curve.) Figure 7 shows lift- curve slope snd maximum lift coefficient as a function of Mach number for the two sections.

The loss of lifwve slope for the NACA OOlC@+ section compared to the NACA 0010 observed in figure 7, approximately.10 percent at Mach numbers below 0.7, cannot be attributed entirely to the differing maximums thickness locations of the two profiles. Reference 4 indicates that some of this deterioration results from the fncreased trail-In&edge angle of the NACA 001&k profile (17O 54* as canpared with 13O 22' for the NACA 0010). The present data do not permit a separate evaluation of the effect of this geometric variable, but the general conclusion is indicated that shifting the maximum thickness to &percent chord decreases the lift-curve slope at all Mach numbers for which data were obtained.

Figure 7 also demonstrates that the -imum lift Coefficient Of the NACA 0010-64 airfoil section, comzpared to the NACA 0010, is appreciably smaller to approximately 0.7 Mach nusiber; but that differences beyond this value are inconsequential. Reference 4, on the other hand, indicates

that the maximum lift coefficient at Mach numbers above 0.7 would be reduced by the increase in the traili~dge angle. By virtue of these facts, it seems evident that some increase of maximum lift coefficient, above 0.7 Mach number, results from the rearwar d shiftofmaximus+ thiclmess location.

DraHivergence Mach number (defFned as the Mach number at which, for a constant angle of attack, the slope of the curve of drag coeffi- cient versus Mach number equals 0.10) is plotted in figure 8 as a function of section lift coefficient. The advantage of the more rearward maximus+ thickness location is here clearly evident, the Mach number for drag divergence being increased about 0.05 at lift coefficients up to 0.4. According to reference 4, however, some of this gain accrues from the change in trailing-edge angle.

Further evidence of the effect of the re arward shift of the maximw~ thickness on drag appears in figure 9, which illustrates the variation of drag coefficient with lift coefficient for the two profiles. A poinl+by;point comparison between the (a> and (b) portions of this figure atMachnumbers abovethatfor drag divergence will quickly demon- strate that the NACA OOlS64 profile has much smaller drag coefficients than does the NACA 0010, but an approximately equal rate of drag rise with increasing lift coefficient. Reference to figure 3 shows that the reduction of drag stems pr3mariI.y froan the delayed drag rise of the NACA OOl.C&4 section as compared to that of the NAC!A 0010 section.

In figwe 10 is seen the variation of pitching-mament coefficient with lift coefficient for the two airfoil sections, and the change tith Mach nw&er of the slopes of these curves at zero lift is illustrated by the plots of figure IL The variation of the slopes displayed by both profiles at the higher Mach numbers is undesirably great, snd it wiU be observed that moving the point of maximum thickness from jC+percent chord to 4&percent chord provides little mrovement.

CONCLUSI~S

A comparison of the experimental lift, drag, and pitching-mament characteristics of the NAC!A 0010 snd 001&b airfoil sections at Mach numbers up to 0.91 provides the foILowIng conclusions relative to changing the maximum-thickness position from 30-percent to @percent chord:

1. The lif+curve slope decreased approximately 10 percent through- out the Mach nuxiber'range of the investigation, the Mach nusiber for lift divergence was practically unaffected, and the maximum lift coefficient was reduced at Mach numbers below 0.70.

2. The drag-dSvergence Mach nur&er increased approxfmately 0.05 at lift coefficients up to 0.4. The rate of drag rise with increasing Mach number above that for drag divergence was virtually unchanged.

.

*

NACA RM 'As19 5

3. The variation with Mach numb&-of the slopes of the pitching- mcunent versus lifticoefficient curves (measured at zero lift) was praw tically unaffected.

Ames Aeronautical Laboratory, National Adtisory Committee for Aeronautfcs,

Moffett Field, Calif.

RElEmNcFls

1. Abbott, -1ra H., van Doenhoff, Albert E., and Stivers, Louis S., Jr.: Summary of Airfoil Data. NACA Rep. 824, 1945.

2. Stack, John, and van Do-off, Albert E,: Tests of l6 Related Airfoils at High Speeds. NACA Rep. 492, 1934.

3. Allen, H. Julian, and Vincenti, Walter G.: Wall Interference in a ~Dlmensfonal-Flow Wind Tunnel, With Consideration of the Effect of Compressibility. NACA Rep. 782, lgkk. .

4. Summers, James L., and Graham, Donald J.: Effects of Systematic Changes of Wailing-Edge Angle and Leading-Edge Radius on the Vmiation With Mach Number of the Aerodynsmic Characteristics of a l&Percent-Chord-Thick NACA Airfoil Section, NACA RM AgG18, 1949.

NACA FM AgGlg

TABLE r.- COORDINATES OF TIIENACAAJRFOILS TESTED

[Stations and ordinates given in percent of airfoil chord]

NACA 0010 SECTIOR NACA 001G-64 SECTION

.

Upper and lower surface

Station Ordinate

0 1.250 i.578 2.500 2.178 5.000 2.962 7.500 3.500

10.000 3.902 15.000 4.455 20.000 4.782 25.000 4.952 30.000 5.002 40.000 4.837 50.000 4.412 60.000 3.803

;IF?% 3.053

go: 000 2.187 1.207

95.000 .672 100.000 .105

L.E. radius, 1.10 percent c

Upper and lower surface 1 Station

0 1.250 2.500 5.000 7.500

10.00-0 15.000 20.000 25.000

OrdLnate

0 1.511 2.044 2.722 3.178 3.533 4.056 4.411 4.666 4.856

E$ 4:433 3.733 2.767 1.556

.856 ,100

L.E. radiu, 1.10 percent c

t ,

NACA 00/O

NACA 00/O- 64

figure I. - Profiles of the NACA 0010 and 00/O-64 airfoil sections.

8 NACA EM AgGlg

s -? CI E I-

O 0: .4 Cl ’ 0 2*

~6 .2 .3 .5 .6 7 .8 .9 I.0

(u) NACA 00/O drfo// section.

fipure 2.-Voriotion of section /ift coefficienf with Mach numhr at vofious ong/es of oftuck.

XACARMAgOlg 9

.8

-.--a I

l-k I’ ffl

Much number, M

ib/ NACA 00/O- 64 oirfo// section.

figure 2. - Condude d.

t

I.

NACA RM AgO

Cl I0 t-t 0 A ;er I/f I I I I Idl I a

I I I t / I t /I

’ I I I I

L I I I I I t I I .4 25 .6 .7 .8 Moth number, M

(0) NACA 00/O oirfoi/ section.

Figure 3.- Voridon of section bog coefficient w/M Mach number of various ongles of otfock

EACARMAgGlg 11

./8 I I I I I I I I l/l I I YI

I ./6.

.

I

v* /

I I I III K&V IPI M I

Much number, M

(b) NACA 00/O-64 uirfoii section.

Figure 3.0 Concluded

I2 XACA RM AS19

I t I I I I I I I

-

- .

-

6 d n

.3 tl .5 .6 .7 Moth number, M

.8 .9 i.0

(0 NAOA OOIO uirfoil section.

Figure 4. - Var idion of section quor ter-cho f d pitchlng - moment Coefficient with Much number of vufious ung/es of 4tt4Ck.

.

ISACA RM AgGlg 13

.3

-337 - I

.4 .S .6 .7 .8 .9 l.0 Mach number, M

(b) NACA 00/O-64 oirfoi/ secfion.

. f iguf e 4.- Conciuded.

. Set

I/i I/I l/l iii iji ijl iii lji iii j ji iii j ii i ji i/i i ji I k’i 1 VI VI VI I/I III III III l/l l/l l/l VI I’I I

-6-1 I ‘I ‘I ’ ’ “I ““““I ’ Jo .40 $0 .55 -60 625 .65 675 .70 225 75 725 &?O .,W &i k??5 $0 Mach number for ~0-00 axis

-4 0 4 8 I2 &on angle of attack, 4, deg ffor M-.3Q

(al NACA 00/O &foil section.

Figure S-Variation of the section lift coefficient with angle of attack of various Mach numbers.

L * . l 1

I , _m

E i 5

I l/l I/I I/I I/I III Ill l/l l/l l/l I/I I/I I/I I/I III G

76’ - ‘. -. ‘. ‘. ’ - ‘. ‘. m ’ m m m ‘. . . . -. - ‘I I I .= .& 050 .55 .60 ~525 65 .675 .?V ,725 .K .M .80 .82Ji .&5 .875 3.3 ’

-4 0 4 8 12 Mach number for % = @“axis

Secflon angle of affucl, a;, deg (for # n.30)

Figure LX- Concluded.

lb) NACA 0010-64 ahfoil section,

c I I AIACA 00/O

s 1 1 ) 1 1 1 1 ---- NAGA (1(//U-64

-4 72 0 .2 .4 6 .8 l.0 12 Section l&Y coefficient, cj

figure 6.- Comparison of the vorhtion of tfte Mocb number for /iff divergence with section liff CO8ffIki8flt for the NACA 00/O and 00/O-64 airfoil sections.

I , . s I

NAC!A RM A%19 17

.8

.6

-- ---NACA 0010-64

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

figufe Z- Compurison ‘of the voriufions of the maximum section liff coefficient and secfion /iff-curve slope with Much number for the NACA 00/O und 00/O-64 uirfoi/ sections.

t I I i I I NACA 00/O I I I I 3 .4

I I I I I I I I I I I I - --- NACA 00/O-64 1 a .==qG&7.

% -45 -4 -2 0 2 4 6 0 Lo I2 section lift coeffktent, 5

Figure &-Comparison of the varialion of the Mach number for drug divergence with sect/on /ift coefficient for Me NACA 00/O and 0010-64 airfoil sections.

IUCAFWAgGl9 . 19

Note: Dotted /he f epresents focus of cd corresponding

.875

.85

.8P5

.80

Section /if# coefficient, c,

(u) NACA 00/O airfoil section. w

figure S.- Vuriution of the secfion drug coefficient wifh /ift coefficient at various Much numbers.

EACA RM AgG19

Note: Dot ted line fepresen ts locus of cd corresponding to h&.

- JO f :s .08 e 9 .06

,

.875

.85

.80

,775

.75

,725

.70

.675

.65

,625

.60

.55

.50

.40

In

I I I I ti 1 i i t 1 .875-1 If

‘, Y I /I -II I I I I /I 1 I/ I/II I I I I

, , ,, , , , I I I I I I.” ’ !A=!!

l ?8 ~6 -.4 ~2 0 .2 .4 .6 .8 40 &2 Section /ift coefficient, c,

-557

(&I NACA 00/O-64 airfoil section.

Figure S.- Concluded.

XACA RM AgGlg 21

d35

.825

&90

-775 d $ -75

El-

7bk

0 Qp725

Section lift coefficient, c,

iid NACA 00/O airfoil section.

Figurs IO.- vufiution of the ssction quu?t8r-chord pitching-momsnt coefficient with lift coefficient at vufious Mach numbers.

22 EACA RM Ag1319

.875

.85

.825

.80

.775

-75

,725 .

.70

.675

.65

.625

.60

.55

30

18 ~6 ~4 ~2 0 .2 .4 6 .8 LO LP Sectton /iff coefficient, cl

(b) NACA 00/O-64 olrfoil section.

Figure IO.- Concluded.

1 I l . . ,

.2 O-

-L/ -

-. P-

-.3-

-. 4-

-. 5-

.3 .4 .5 .6 .I .8 .9 -- -- 1 -- ---

- NACA 00/O

---- NAGA 00/O-64

L 0

Figure IL- Comparison of the voriotion with Mach number of the rute of change of section quo&w-chord pitching-moment coefficient with lift coefficient, of the N.&A 00/O and 0010-64 &foil sections.