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RESEARCH MEMORANDUlv--iJFT, DRAG, AND PtiCHING MOMENT OF LOW-ASPECT=RATIO WINGS

AT SUBSONIC AND SUPERSONIC SPEEDS - PLdVYE

TRIANGULAR WING OF ASPECT RATIO 4

WITH NACA 0005-63 SECTION

By John C.. Heitmeyer and Jack D. Stephenson

By....... ...........z.g...@57..57

]k$%.................................. ,,,.........................------.....GRADE OF Oi FICL:. ~~1 b LHAkGE)

NATIONAL ADVISORY COMMITTEEFOR AERONAUTICS

WASHINGTONFebruary 2Y 1951

. . . i-,.,!v-:...-- ‘

— — —

1.

.

NACA RM A50K24

TECH LIBRARY K/WE, NM

I’IIIIIIIIIIIIIIIIIJ:IIIIIIIIIIII‘“-”“=-=nlJ42745 : “:.-

NATIONAL ADVISORY,COMMITTEE FOR AERONAUTICS

RESEARCH MEMORANDUM●

LIFI, DRAG, AND PITCHING MOMENT OF IOW-ASPEC!I+ATIOWINGS

AT SUBSONIC AND SUPERSONIC SPEEDS - PLANE

TRIANGULAR WING OF ASPECT RATIO 4

WITH NACA 0005-63 SECTION

By John C. Heitmeyer and Jack D. Stephenson

SUMMARY -

A wi~-body conibination having a plane triangularratio & and NACA 0005-63 sections in streamwise planestigated at both stisonic amd supersonic Mach numbers.

wing of aspecthas been inves-The lift, drag,

and pitch- moment of the model are presented for Mach numbers from ‘0.25 to 0.g6 snd 1.20 to 1.70 at a Reynolds number of 1.5 millign. Thevariations of the characteristicswith Reynolds nunher are also shownfor several Mach numbers.

INTRODUCTION

A research program is in progress at the Ames -AeronauticalLabora-tory to ascertain expertientally at subsonic and supersonic Mach numbers

—

the characteristics of wings of interest in the design of hig&speedfighter airplanes. Variations in plan form, twist, ember, and thick-ness are being investigated. This report is one of a series pertainingto this program and presents results of tests of a -body conibinationhaving a plane triangular wing of aspect ratio 4 and NACAOOO%3sections in streamwise planes. Results of other investigations in thisprogram are presented in references 1 and 2. 4s in these references,the data hereti are presented without analysis to expedite publication.

NOTATION

b wing spsa, feet

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2 NACA RM A50K24.

()b/2 ~2 dyL

mean aerodynamic chord ,feet .

$/2 c dylocal wing chord, feet

length of body includ~ portion removed to accommodate sting,inches

lift-drag ratio

maximum lift-drag ratio

Mach number

fre+stream dynsmic pressure, pounds per square foot

Reynolds num”er based on the mean aerodynamic chord

radius of body, inches

maximum body radius, inches

.

.

- total wing area, including area formed by extending leadingand trailing edges to plane of symmetry, square’feet

longitudinal distance.fromnose.of body, tithes

distance perpendicular to plsme Of symmetry, feet

angle of attack of body axis, degrees

drag coefficient ‘dxag

0 qs

lift coefficient.()

liftT

pitcheoment coefficient referred to

aerodynamic chord(

pitching moment

qS5 )

slope of the lift curve measured at zero

.

—

quarter point of mean

lift, per degree

slope of the pitchi~oment curve measured at zero lift-.

h

—

NACA RM A501C24 3

APPAEMTUS

Wind Tunnel and Equipment

The experimental investigateion was conducted in the Ames I-2-footpressure wind tunnel ad in the bes 6 by &foot supersonic wind tunnel.In each wind tunnel the Mach number can be varied continuously and thestagnation pressure can be regulated to maintain a given test Reynoldsnuriber. The air in these tunnels is dried to prevent formation of CO*densation shocks. Further information on these wtid tunnels is pr+sented in references 3 and 4.

The model was st~ mounted in each tunnel, the diameter of thesting being about 82 percent of the diameter of the body %ase. Thepitch plane of the model support was vertical in the Wfoot wind tunneland horizontal in the & by &foot wind tunnel. A balance mounted onthe sting support and enclosed within the body of the model was used to.measure the aerodynamic forces and moments on the model. The balancewas a &l/2-inch, fo~omponent, stra~age balance of the t~edescribed in reference 5..

Model

A photograph of the model mounted in the Ames M!-foot pressure windtunnel is shown in figure 1. A plan view of the model and certain modeldimensions are given in figure 2. Other important geometric characte~istics of the model are as follows:

wing

Aspect ratio . . . . . . . . . . . . . . . . . . ...4Taper ratio . . . . . . . . . . . . . . . . . . . . ..0Airfoil section (streamwise] . . . . . . . . NACA 0005-63Totalarea, S, square feet . . . . . . . . . . . . 2.007Mean aerodynamic chord, 6, feet . . . . . . . ,. . . 0.944Dihedral,degrees. . . . . . . . . . . . . . . , . . 0Camber . . . . . . . . . . . . . . . . . . . . . . .NoneTwist,degrees . . . . . . . . . . . . . . . . . . . . 0Incidence, degrees . . . . . . . . . . . . . . . . . . 0Distance, wing+hord plane to body axis, feet . . . . . 0

NACA RM I@K24.

Body-_

Fineness ratio (based upon length Z; fig. 2) . . . . . 12.5Cross-section shape . . . . . . . . . . . . . . . . CircularMS,Ximumcross-sectional area, square feet . . . . . . 0.1026Ratio of maximum cross-sectionalarea_& wing area . . 0.0509

The wing was constructed of solid steel. The body sp~ was alsosteel and covered with alumlnum to form the body contours. The surfacesof the wing and.body were polished smooth. .—

TESTS AND PROCEDURE

Range of Test Variables

The characteristics of the model (as a function of angle of attack)were investigated for a range of Mach nwibers from 0.25 to 0.96 in the *.:Ames 12-foot pressure wind tunnel and from 0.60 to 0.93 and from 1.20 ..to 1.70 in theportion of theData were alsonumber of 0.25Mach numbers-.

Ames 6- by 6-foot supersonic wind tunnel. The majordata was obtained at a Reynolds number of 1.5 million.

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obtafned for Reynolds nunibersup to 8.o million at a Machand up to a Reynolds number of 3.0 million at supersonic

Reduction of Data

The test data have been reduced to standard NACA coefficient form.Factors which-could affect the accuracy of these results and thecorrections applied are discussed in the following paragraphs.

Tunnel-wall interferehce.- Corrections to the subsonic results forthe induced effects of the tumnel walls resulting from lift on the modelwere made according to the methods of reference 6. The numerical valuesof these corrections (which were added to the uncorrected data) were,for the results from the 12-foot wind tunnel:

--

&= 0.14 cL

&!D = o.oops cL2

.

—.—

.

NACA RM A50K24.

5

and, for the results from the & by &foot windd

AZ= 0.47 CL

ND = 0.0081 %2

No correci.ions were made to the pitching+mmnt

The effects of constriction of the flow at

tunnel:

coefficients.

subsonic speeds by thetunnel walls were taken into account by the mezLod of reference 7. Thiscorrection was calculated for conditions at zero sngle of attack and wasapplied throughout the amgle-of+ttack ramge. Ai aMachntier of 0.96in the 1.2-footwind tunnel, this correction’amounted to a l-percentincrease h the Mach number over that determined from a calibration ofthe wind tunnel without a model in place. In the 6-by &foot windtunnel at a Mach nuuiberof 0.93, the similar correction was 3 percent.

For the tests at supersonic speeds, the reflection from the tunnelwalls of the Mach wave originating at the nose of the body did not crossthe model. No corrections were required, therefore, for tunnel+alleffects.

.Stream variations.- Calibration of the U&foot wind tunnel has

shown that in the test region the stream inclination determined fromtests of a wing spanning the tunnel, with the support system at 0° angleof attack, is less than O.O&. The vsriation of static pressure is lessthan 0.2 percent of the dynamic pressure. No correction for the effectof these stresm variations was made.

Tests at subsonic speeds in the &by 6-foot supersonic wind tunnelof the present symmetrical rmdel in both the normal and the invertedpositions have indicatedno stream curvature or inclination in the pitchplane of the model. No measurements have been made, however, o’fthestream curvature in the yaw plane. At subsonic speeds, the longitudinalvariation of static pressure in the region of the model is not knownaccurately at present, but a preliminary survey has indicated that it isless than 2 percent of the dynwnic pressure. No correction for thiseffect was made.

A survey of the air stream in the &by 6-foot wind tunnel at supe~sonic speeds (reference 4) I&! shown a stresm curvature only in the yawplane of the model. The effects of this curvature on the measured chapacteristics of the present model are not lmown, but sre believed to besmall as @lged by the results of reference 8. The survey also indicatedthat there is a static=pressure variation in the test section of suffi-cient magnitude to affect the drag results. A correction was added to

. the measured drag coefficient, therefore, to account for the longitudinalbuoyancy caused by this statibpressure variation. This correction

*

6- NACA RM A50K24.

varied from as much as -0.0016 at a Mach number of 1.20 to +0.0016 at aMach number of 1.70.’ .

Support interference.- At subsonic speeds, the effects of supportinterference on the aerodynamic characteristics of the model are notlmown. For the present tailless model, it is believed that such effectsconsisted primarily “ofa change in the pressure at the base of the model.In an effort to correct at least partially for this support interference,the base pressure was measured and the drag data were adjusted to correspend to a base pressure equal to the static pressure of the free stream. —

At supersonic speeds, the effects of support interference of abody-sting configuration similar to that of the present model are shownby reference 9 to be confined to a chemge in base pressure. The pre- 4 :viously mentioned adjustment of the drag for base pressure, therefore,was a~lied at supersonic speeds.

Errors introduced by support system.- Clearances between movingparts in the support system in the 6- by &foot supersonic wind tunnel .

under certain conditions permitted the angle of attack to vary as muchas 0.3° with no c-e in the angl~f-attack indicator. The clearanceswere discovered after inspection of the data herein showed that the drag

.—

coefficients were not the seineat positive and negative lift coeffi-cients. However, calibration of the angle-of~ttack indicator had beenmade in such a manner that the angles of attack and thus the lift anddrag results were correct at positive lift coefficients. Further proofof this fact was”obtained from reruns at several Mach numbers made In amanner to eliminate altogether the effects of the excessive clearsnce.The drag data from these tests (symmetrical.about zero lift) ~eed withthose of the former tests at Eositive lift coefficient, as did the angleof attack and lift and pitch~oment coefficients.

lanee.- As the model is pitched in the vertical plane in thel>foot wind tunnel, the weight of the model produces a change in themeasured forces and moments, which for the present tests was significantonly for the chord-force measurements. The measured chord-fone tarehad a small discontinuity when the chord force reversed direction.Since the same discontinuity was present inthese data were corrected for this inherenturing system.

RIHJErs

the uncorrected drag data,characteristic of the meas-

!Theresults are presented in this report without malysis in orderto expedite publication. Figure 3 shows the variation d lift coeffi-cient with angle of attack and the variation of drag coefficient,

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NAC!ARM A50K2k t’”=””~ 7

pitchi~oment coefficient, and lifti=ag ratio with lift coefficient. at a Reynolds number of 1.5 million smd at Mach numbers from 0.25 to

1.70. The effect of Reynolds number on the aerodynamic characteristicsat Mach numbers of 0.25, 1.20, and 1.53 is shown in figure 4. Theresults presented in figure 3 have been summarized in figure 5 to showsome important parameters as functions of Mach number. The slope par-eters in this figure have been measured at zero lift.

Ames Aeronautical Laboratory,National Advisory Committee for Aeronautics,

Moffett Field, Calif.

REFERENCE

1. Smith, Donald W.; and Heitmeyer, John C.: Lift, Drag, and PitchingMoment of LmeAspec~atio Wings at Subsonic and Supersonic.Speeds - Plane Triangular Wing of Aspect Ratio 2 With NACAOO08-63Section. NACARM A50K20, lg50.

2. Smith, Donald W., and Heitmeyer, John C.: Lift, Drag, and PitchingMoment of Iow-Aspec~atio Wings at Subsonic and SupersonicSpeeds - Plane Triangular Wing of Aspect Ratio 2With NACA COO%3Section. NACARMA501Q1, 1950.

3. Edwards, George G., and Stephenson, Jack D.: Tests of a TriangularWing of Aspect Ratio 2 in the Ames 12-Foot Pressure Wind Tunnel.I - The Effect of Reynolds Number and Mach Ninnberon the Aer-dynsmic Characteristics of the Wing with Flap Uhdeflected.NACARMA7K05, 1947.

4. Frick, Charles,W., -d Olson, Robert N.: Flow Studies titheAsymmetric Adjustable Nozzle of the Ames & by &Foot E@ersonicWind Tunnel. NACARMA9E24, 1949.

5* Olson, Robert N., and Mead, Merrill H.: Aerod-c Study of aWin@?uselage Combination @loying a Wing Sw@ Back 630.–Effectiveness of an Eleven as a Imgitudinal Control and theEffects of Camber and Twist on the Meximum Lift-Drag Ratio atSupersonic Speeds. NACARMA50A31a, lg50. ,

6. Glauert, H.: The Elements of Aerofofl end Airscrew Theory. TheUniversity Press, Canibridge,England, 1926, ch. XIV.

8

7. Herriot, John G.: Blockage Corrections for Three+Dimensional-FlowClosed+!hroat Wind Tunnels, with Consideration of the Effect ofCompressibility. NACARMA7B28, 1947.

._

.

8. k6Sh& Henry C.: Aerodynamic Study of a Win&Fuselage Combinationlhnployinga Wing Swept Back 63°- Effect of Sideslip on Aer*dynsmic Characteristics at a Mach Number of 1.4With the WingTwisted and Csmber?d. NACARM A50F09, 1950.

9. Perkins, Edwrd W.: ~perimental Investigation of the Effects ofSupport Interference on the Drag of Wdies of Revolution at a MachNumber of 1.5. NACARMA8B05, 1948.

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*# ,

Equotion of fweluge radii

2X2%&=[f-(l-i) ]

All dimenff’ons shown in inches

unless ofherwise noted

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I 45.38 B

Figure 2.– Front and plan views of the modeL

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10

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+5

+ ‘4 O 4 8 E I(5 zo for M= .g5

Angle of ohbck, a, cieg

(a) c--Vs cr.

Fl@n? 3.- 7iie var@ibn d the a&a@kWc chamctwiefics Wflh W tiffcient at Wrh?w Mach numbers.

Rqvna.Us mmkw, 1.5 mdibn.,.

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d I I I I I— “w-.

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L u ~ o Re -“ - - “-’’”’-

1❑ Riwffs fFotn Ames 6’x6’W! Z

76.04 0 -.04 -.08 -12 -.16 -.,2V for Itf.,25

-wi5i=Pitchg-mawent coa?%hmf, ~

/=i@193.- Ca#z&

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1.2

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m 1 1 1 w ,Iii Ifi

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o .04 .06 .12 ./6 .20 .24 .28 .32 fw M= 27

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0 .1 ,2 ,3 .4 for iW=.25

Lift coeffiint, CL

R@?i!? 3.- CMIwded.

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w “4

g’ .2

t o

5-.2

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-.60 .04 .08 .12 “.16 -8 -4 0 4 8 :, 12. :>04 ,-? 0 -.~

bog coefficient, ~ Ar@e of aifockz C, deg PM%7g-motmwt CMl%t%%t, Cm

(a) hf=.25.

Figure 4.- The m“btiw of the otwdwk ,Ghoractiistics

lift coef17cie$7f at various Reynolds numks

,, 1

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,4@

$,.2

Q$

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c+ “6.

1. .4\

1 .2Q

ao

-.2

-“40 .04 .08 .12 .16 .20 -4 0 4 8 L? .04 0 -.04 -.08 -.12 -.)6 -.20

Dug &7eft7ci?rJt, ~ At@ d attack, cc, tfeg Plwahg-nwnenf CaXw%w, ~

=is=’(cl M= I.5.3.

d. 1, .,, f,

NACA RM A50K24 19

●

o .2

:3\

1

:2 : ++/

I

—

C$(&h~ N~ 0005-63

*IW wing

-J - / >/

“- — Resu/ts from Ames /2’ W Z

— — Results from Ames 6’x 61 H! T 1

n ~“o .2 .4 .6 .8

Much number, MdC*

(b)aq Vs M-

chowcterktibs as a function ofFigwe 5.- Summary of uwodynomic

Mch number. Reynolds number, /.5 million..

20

24

20

8

4

n

M-NACA RM A50K24

\ . .

w -. ~ -\ ,

. .

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-+ +-

NACA 0005-63

Plane wing

Results from Ames /2’M Z—— . Results from Ames 626’ U!Z

.

-,

.

~—–.o .2 .4 .6 ..8 /.0

Mach number, M

.4

Qjj

./

/.2 /.4 /.6 /.8.

M —

.2 “ .4 .6 .8 /.0 1.2 /.4 1.6 1.8Mach number, M

v—. ...- .

(d) (/)c’ fOr ‘/3 max Vs M..

Figt.re 5.- Con#wed.

b

NACA RM A50K24 . 21

.16

./4

./2

@./0

+

.~

.Qg .08

8

~ .06

.04

.02

0

— Results from Ames 12 t W Z“+ ‘ -A{‘“ ‘~

/

— —Results from Ames 6x6’ W.Z ,/AMCA 0005-63

Pbe wing

~

/ “- - ~ .

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- .6 /

0“/ “

4~0“

//

/ — —— > - .4-/

/

-.2 / —

/ “ o=s=’

o .2 .4 .6 .8 10 12 !4 L6 L8

Mach number, M

(e} CD vs M.

fi~e 5.- Concluded

NACA-UI@ey -2-2-51- 92]