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SECURITY INFORMATIONrL “t Cmv e9*

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1

RESEARCH MEMORANDUM”-”””LIFT, DRAG, AND PITCHING MOMENT OF LOW-ASPECT-RATIO WINGS

AT SUBSONIC AND’ SUPERSONIC SPEEDS - PLANE TRIANGULAR

WING OF ASPECT RATIO 3 WITH AIR-TO-AIR

MISSILE MODELS MOUNTED EXTERNALLY

By Donald Conrard

Ames Aeronautical LaboratoryMoffett Field. Calif.

Pd’-JJ7F/6aClassification cancelled (or changed to .... .......... ........................)

BY Authority of l~z~ ~~<k.’~~kdqti~~.?~.... . ....... ...(OFFICER AUTHORIZED TO CHAWE)

By. .........--.-ii..E ~~iD afgk+-z-----...............................,.. ,‘- ................................................. ...........................

GRADE OF OFFICER MAKING CHANCE)

NATIONAL ADVISORY COMMITTEEFOR AERONAUTICS

WASHINGTONJune 26, 1952

. . .

TECH LIBRARY ~FB, NM .— ..—

l!lll!Hlll[ll[lll!lll!lll[llllfgNACA RM A~2C10a alJ42a74— -. .—

NATIONAL JKOVISORYCOMMITTEE FOR

RESEARCH MEMORANDUM

AERONAUTICS

LIFT, DRAG, AND PITCHING MOMENT OF LOW-ASPECT-RATIO WINGS AT

SUBSONIC AND SUPERSONIC SPEEDS - PLANE TRM~

WING OF ASPECT RATIO 3 WITH AIR-TO-AIR

MISSILE MI)DELSMOUNTEDEXTERNALLY-

By Donald Conrard *>

SUMMARY

This report presents results of an investigation of effects ofexternally mounted missile models on the aerodynamic characteristicsof a triangular wing of aspect ratio 3 at both subsonic and supersonicMach numbers. The lift, drag, and pitching moment of the ylng-fuselage

# model fitted with the missile models are presented for Mach numbers of0.6, 0.9, 1.2, 1.4, snd 1.7 at a Reynolds number of 4.8 million.Similar data me presented for the basic wing.

w.

INTRODUCTION

A research program is in-progress at the Ames Aeronautical.Laboratory to ascertain e~ertientally at subsonic smd supersonicMach numbers the characteristics of wings of interest in the desiaof high-speed fighter airplanes. The present report is concernedwith the influence of externtiy mounted missiles on the characteristicsof a wing-body combination incorporating a plane triangular wing ofaspect ratio 3. The model is the same as that used in reference 1.The present investigation was Mmited to tests of the basic wtig tidthe basic wing fitted with six models of an air-to-air missile of thebesm-rider type and, alternately, eight models of an air-to-airmissile of a passive-seeker type.

As in reference 1, the data here; sre presented without analysis.“ to expedite publication.

v

I

2. NACA RM A52C1OEL

b

NOTATION

wing span

()

~-i= ,2@ ‘

meau aerodynamic chord —

J~/2 Cdy .0

c local wingchord.7

1 length of body including portion removed to accommodatesting

Llift-di+agratio

5

()

& .’ max- lift-dragD m=

M.

q

R

r

ro-

S

x

Y

a

CD

ACD

CL

Mach number

--.

ratio

.-

free-stresm dynsmic pressure ‘-

Reynolds number based on the meaz-aerodyusmic chord..

!+

&

-L

.

-. -:

.

. ..=

.-

—. : -. .<--

- .-

radius of body

maximum body radius

total wing area, including area formed by extending leating._and trailing edges to plane of symmetry

longitudinal distance from nose of-body

distance perpendicular to plane of symmetry

angle of attack of body =is, de~ees

drag coefficient()‘drag,F

incremental drag coefficientmodels based on wing area

lift coefficient()

liftF

of externally stored missiles

—.

.-..

.-. :

.

u

t

NACA RM A52C108

* cm pitching-moment coefficient

(

referred to quarter p’ointof

mean aerodymmic. chordpitching moment

n qse )

dc~

zslope of the lift curve measured at zero lift, per degree

dCm

qslope of the pitching-moment curve measured at zero lift

APPARATUS

Wind Tunnel and Equipment

The ecqertientsl.investigation was conducted in the Ames6- by 6-foot supersonic wind tunnel. In this wind tunnel, the Machnumber csm be vsried continuously and the stagnation pressure can beregulated to maintain a given test Reynolds number. The air is driedto prevent formation of condensation shocks. Further information onthis wind tunnel is presented in reference 2. ..

The model was sting mounted in the tunnel, the diameter of thesting being about 93 percent of the diameter of the body base. The .,

* pitch plane of the model support was horizontal. A A-inch-diameter,four-component, strain-gage

9 enclosed within the body ofdynamic forces and moments.

balance (described in reference 3),the model, was used to measure the aero-

Model

A photograph of the model is given in figwe 1. Figures 2, 3,4, and ~ give important dimensions of the basic wing, the missilemodels, and the arrangement of the missile models on the triangularwing.

4

wing

Aspect ratioTaper ratio

NACA RM A52C10a ‘--”-“.

.—~.

● ***O** ● ***O.* ● .***** ● **9.. 3“ ‘“ r:. .... *..* . . . . . . ● .* .* . .. _._. . . . . . 0Airfoil section (streamwise) . . . . . . . ..Total area, S, square feet . . .“.“~“. . . .Mean aerodynamic chord, Z,fqet. .:-. . . ;-Dihedral, degrees . . . . . . . . . . i . . .Camber. . . . . . . . . . . . . . . . . . . ..!l?wist,degre ea.. . . . . . . . . . . . . . .Incidence, tiegrees. . . . ... . . . . . . . .Distance, wing-chord plane to body axis, feet

.*. NACA 0003-63

. . . . . . . ● 2.k25 .“- ,

.****. -–9,1499” ,

..**** . . . 0 .-. .:● *...* . None . , .:...>. ● .*0** ● *O

– :;.__. ....:’. ...**. ● 9... *..*. . . 0

Body.,..- ..

Fineness ratio (based upon length Z;-”fig.2) . ● ● ● . ●“ ~“ 1-2:5 ._.Cross-section shape . . . . . . . . . .-. ‘. . . . . . . C!i”rcu@ . ----

Maximum cross-sectional area, square feet . . . . . . . . 0.1235 - .:Ratio of maximum cross-sectional area to wing area.. . . . 0.0509

Thefuselagethe bodysmooth.

The

The

-..model wing was constructed”of solid steel ahd the modei

..27’=”.:;. ..m-c-

onsisted of a steel spar covered-with aluminum which formed” ~- ...contours; the surface “ofthe WQ and b&iy were polished ““ ““””““,,. -:”-.—.=-

missile models were constructed entirely “ofst”eel. ““. .

*..—.”3

TESTS AND PROCEQURE .,. ,“.-

Range of Test V&riables—. .--. .—

.-, .:- :

characteristics of the model (as “afunction of sngle of.

attack) were investigated for Mach numbers of 0.6, 0.9., 1.2, 1.4, and .- .1.7. The data were obtained at”a constant Reynolds numbek of4.8 million.

_

Reduction of Data

,, , L_.=

The test data have been reduced to sttidard NACA coefficient _.form. Factors which could affect the accwacy of these results,

.

together.withpar~aphs.

the corrections applied, are..discussedin the foliowing.6

w

NACA A52c10a ~ 5

Tunnel-wall interference.- Corrections to the subsonic resultsfor the tiduced effects of the tunnel walls resulting from liftmodel were made according to the methods of reference 4. Thenumerical values of these corrections (which were added to theuncorrected data) were:

on the

.. .. .

La = 0.554 ~

ACD = 0.0097 cL2

No corrections were made to the pitching-moment coefficients:

by the5= - -.

The effects of constriction of the flow at subsonic speedstunnel walls were taken into account by the method of referenceThis correction was calculated for conditions at zero angle of attackand was applied throughout the angle-of-attack range.

.._At a Mach

number of 0.90, this correction amounted to a.2-percent increase inthe Mach number and in the dynsmic presstie over that determined froma calibration of the wtid tunnel without a model in place.

For the tests at supersonic speeds, the reflection from thetunnel walls of the Mach wave originating at the nose of the body did .

..

not cross the model. No corrections were required, therefore, fortunnel-wall effects.

Stresm variations.- Tests at subsonic speeds in the 6- by 6-foot --supersonic wind tunnel of the present symmetrical model in both the .normal and inverted positions have indicated a stresm inclination of-0.05° amd a stream ctivature capable of producing a pitching-moment

-.-——

coefficient of -O.(M4 at zero lift. No corrections were made to thedata of the present report for the effect of these stream irregulari- ~ ;ties. No measurements have been made of the stream curvature in theyaw plane. At subsonic speeds, the longitudinal variation of staticpressure in the region of the model is not known accurately atpresent, but a preliminary survey has indicated that it is less than2 percent of the dynamic pressure. No correction for this effect wasmade. —

A survey of the air stream in the6- by 6-foot wind tunnel atsupersonic speeds (reference 2) has shown a stresm curvature only intheyaw plane of the model. The effects of this curvature on themeasured characteristics of the present mcdel are not Mown but are ‘

. ..-

believed to be small as judged by the results of reference 6. The.-

survey of reference 2 also indicated that there is a static-pressurevariation in the test section of sufficient magnitude to affect thedrag results. A correction was added to the measured drag .-

.

6 } NACA RM A52C10a

coefficient,‘therefore,to account for the.longitudinalbuoyancycaused by this static-pressurevariation. .Thiscorrect+onvariedfrom 0.0002 at a Mach number of 1.2 to 0.C@6 at a Mach number of 1.7

Support interference.- At subsonic speed:,the effects of supportinterference on the aerodynamic.characteristics of the model are notknown. For the present tailless modeI,-it-~s believed that sucheffects consisted primarily of a “changein the pressure at the base ofthe model. In an effort to correct at least partiaJJy for this

.*“*

-... .—,e

_.

..-..-.- -:-

.. .. . .=

.- *—.—.

support interference, the base pressure was-measured &d the drag data.were adjusted to correspond 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 theshown by reference-7 to be conftied to a ch~gepreviously mentioned adjustment of the drag fortherefore, was applied at supersonic speeds.

RESULTS” ~

present model-are .in base pressure. The “base pres~ure,

The results are presented in this report without an_glysisinorder to expedite publication. The variation of lift coefficient withangle of attack and the variation of pitching-moment coefficient, dragcoefficient, and lift-drag ratio with lift cclefficientqt..Machnumbers ‘from 0.6 #o 1.7 and at a Rejnolds number of”4.8”mill.ionare given in ..figures 6, 7, and 8. The data in figure 6 are those obtained with thelarge air-to-air missile models (beam-rider type) mounted in anunbanked position below the model wing in boh the fore and aft

-_ .-..-

positions-on the model wing. Figure 7 preseRts similar-results for the ...—--large missile model banked 45° when-positioned below the wing. —

Figure 8 shows similar data for the small air-to-air missile for the twochordwise positions on the wing.

The incremental drag coefficients for the two missile models inthe two chordwise positions for two lift coefficients =e given infigures 9 and 10, “=d the maximum lift-&ag~~~tiOaS a .tiction ofMach number is given for all the arrangements in figure Il.

.

In all cases, the aerodynamic characteristics of the basic wingare presented for comparison where appropriate

Ames Aeronautical LaboratoryNational Advisory Committee

Moffett Field, Calif.

.-

. .

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m..”./

~..-:.

. -.

.,.

+for Aeronautics

.-..,

.

.

NACARM A52C10a 7

RIWERENCES

1. Eeitmeyer, John C.: Lift, Drag, snd Pitching Moment ofLow-Aspect-Ratio Wings at Subsonic and Supersonic Speeds - Plane ‘:”Triangular Wing of Aspect Ratio 3 With NACA 0003-63 Section.NACARM A51H02, 1951’.

.“:

2. Frick, Charles W., and Olson, Robert N.: Flow Studies in theAsymmetric Adjustable Nozzle of the Ames 6- by 6-FootSupersonic Wind Tunnel. NACARMA9E24, 1949. .-

3* Olson, Robert N., md Mead, Merrill H.: Aerodynamic Study of aWing-Fuselage Combination Employing a Wing Swept Back63°.- Effectiveness of an Eleven as a Longitudinal Control and “- “’the Effects of Csmber snd Twist on the Maximum Lift-Drag Ratioat Supersonic Speeds. NACARM A~A1.3a, 1950.

.

4. Glauert, H.: Wigd Tunnel Interference on Wings, Bodies andAirscrews. R.& M. No. 1566, British A.R.C., 1933.

.—.—

5* Eerriot, John G.: Blockage Corrections for Three-D@ensionalQFlowClosed-Throat Wind Tunnels, with Consideration of the Effect 6f ---- -Compressibility. NACA Rep. 995, 1950. (FmmerlyN ACARMA’7B28) —

6. Lessing, Henry C.: Aerodynamic Study of a Wing-FuselageCombination Bnploying a Wing Swept Back 63° - Effect of Sideslip “m,-on Aerodynamic Characteristics at a Mach Number of 1.4 With theWing Twisted and Cambered. NACA RM A50F09, 1950.

.-

7* Perkins, EdwardW.: Experimental Investigation of the Effects ofSupport Interference on the Drag of Bodies of Revolution at aMach Number Of 1.5. NACA~ 2292, 1951.

..-.

—

.-.

-.

,.

——

-=

-..

. —-.

—.

.

. ..

K’ NACA RM A52C10a

..

9

Figure l.- Photograph of the model mounted in the “testsection ofthe wind tunnel with small missiles in the forward position.

....

.

---(

~QUUtiM of fU8dO~ radii

@-w”4

-—

P,0

--.=.

r /Z32 — -1’--’8--’1● 46.93 —

-----—..----

● )859.50 I

Figure .?. - Hant ondpkw views &the mo& ting and he@e.

a’

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“d--+2.357 J

i!$!l.674

+–

(227

+“.032

%-

l--—--’-l’”’”-l t-’-l

Large missik

““’ -+=-+Al’ ohemkms in imhes ,Z ,=,m ~g_

f

-q p- ,091I

Figure 3. - Ai~boir missile models utifized h the inwh@otlm.

.

Lcrge m&le

A# dimensions in inches ~

.08

1’

f r-%===iSectbn A -A

Sm#ni$sk

Figure 4.- Method of mounting mksi% models.

I

,,

1 ,, ,,r,>

L >

1

-6 -+ -2 0 .2 4 6’8 ~ EM A$ fo? M-,mo

AI@ d attack, a, dag

(d CL6 a

F~e 6.- Aer@mc cM*- ~ #e trjo~ &gnwzfe/wfM ondwh%tt hr@ fmWe modek mnmted

extemofly In two chodwi~ positions. M%s&sI$wI&$ P. t?eytbkts numb% 4.8 m/Iion.

r4=

.

NACA RM A52C10a

.

.

15

,

.

.8

.7

.6

5

.4

o

-./

-.2

-.3

-.4

.08 .04 0 -m -05 --E tir M=a50

Pitching-moment coefficient Cm

.

(b) CL vs Cm

Figuns 6. -Conked.

I

I I 1- 1 I I I I- i

m

I*I 1/471 /i”/f I I/w I 1/%’7 1111

I I I I I I i I I I I I I I I I I I I I I I I I I I I I I I I I

off P.03u41M OsQ7ut,i29

bag coeffkfem’, ~

(c) c~ VSG

17gwe 6. -continued.

1

3 ,

4.,

!!,.+

,,,..

,

,, .,,, FL, ,r.r. , ,1~I

-.

( .,

.s

4

0

I0 .1 -2.3 .4 .5 .6 7 .8 b M= 0.60

Lift coeffkient, CL

(d)G@

Fgwe 6.- Concllnfed. 5

.

,1

[

.7

6

5

.4

6’

0

-J

-2

-.3

0

El

-4-++l,,,,,l,m,,l,,,,,

I 1 1 1 J 1 I 1 I /1 1 1 M 1 1 1 w 1 1 1 1 1A 1 1 1 I ,4 1 I 1 1

1 /A I P !.4/ 1

I I 1

1 1 1 1 I 1 1 1 1

I

I 1 I I

I I I=QS=-

1 I 1 1 1 WI I 1 I I I I I I I 1 I I

-.4

-6 -4 -P O P 4 6 8 JV L? 14 M * A&w

Anglsc+faffack, a,deg

(a) CL is a

/7@uez - Aelodmk c#m?cfenSJ%S of fbeikngt4r W@ model wifh m?dwh%ouf large mk.sik models mounfwi exftwwb

in two ck?tvbxse positions. Al&si& banked 45”. Reynab!s num$er.4.8 mllllon.

Pa

( ,\ . I I

.7

.6

.6

.4

.3

6’

-.I

-.2

-.4

.Lw 04 o#-LW-J2 -16 fir M =0.60

PiWLOlg-mW??W W9ffk%Wl~ t%

(b) CL w Cm

f@ui5 7.-Gmbh&ti

I

—.

B

.7

.6

.3

o

-./

-.2

-.3

-.4

0 a Lwm.min.txmmm .m ./1 .1. h Aktklv

&ag tawfftimt q)

(c) Q w%

F&m? 7 .-cw?bhud

,,,, ,

lo0

# . # . * 1

20

18

16

14

-b

.$ ‘2P

6

4

2

o’

-+

-=3-++

iI .

,, I , ,, ,,,

III,,

El NI I I I

X-1’ 0.9 1 1.2 / [4 / /.7

I I r---

I 1 r 1 I r 1 I F 1 1 1 ! 1 1 1 1

0 J .2 .3 .4 .5 .6 .7 .8 b h+= 0.60

Ljft coefficient, CL

(d) C&j

/%z+w 7 .-cam$mbd

I

.—

.8

.7

.6

.5

.4

0

-. I

-.2

-.3

. .-6-4 -2 ~ P 4 6 8 IO 12 14 16 for M-O. 60

Awgle of attack, a, deg

(a) Q W a

Ffgwe 8.-Aero@o@ ChOMCWMCS of the triu@r W~ model W#h an$wifh@t smoff &sHe models mtwted

extemu@ in tm akvd~e ~itiw. Rw&& mmber. 4.8 mL&.z

Ivm

{, I .

-./

-. .?

‘.3 H—H—w—H-4 ILuLLLu“:W .04 0 -.04 -.W -.12 -.16 for M=O.60

Pifdhg-mw?hwt coeff~~ &

(b) ~ w Cm

F&e 8.-Cm%zMi.

. .-1

1“

.8

.7

.6

.5

.4

&

5 .,

0

71

-.4?

-d.7

0 :01 .02 .6!3 .W .m .06 .07 .08 .09 JO .11 A? fw Af=O.60

L%zlcaeffkient ~

(c) Q Vscg

-8 .-~

.,,, 1: ‘“ .1

,; ,,

t > .1’

,. #b

1 I i., :,’

II

E!

I

t

# ,

I

20

18

16

14

6

4

2

0

0 .i .2 .3 # .5 .6 .7 .8 for M=O.60

~/ft dfk?ient, CL

.-

26 NACA RM A52c10a i=

.4 .6 .8 [0< /2 14 L6 18

hfach number, 4/,

.010

.4 J$ .8 W /2 14 /.6 18Ak@ number, M’

.

. ,.-

. .—

..

Figure 9.- Incremental drug coefhids fix the missile models” stored,-

externaly as a function of Mach number at zero /ift..

Reynolds number, 48 miltin. --

●

—

.

NACA F/MA52i10a 2’7

.

.

.

.

.

.m---4-- M forward

-V- M aft

.

.mi

.4 .6. .8 10 12 14 16 18

Mach number, M

.0/0

0.4 .6 .8 LO L2 L4 i6 i8

Mach number, M

Figure 10. - I.crementol drag coefficients for the missi/e modelsstwed externally os o function of A&at number at o liftcoefficient of 0.3. Reynokfs number, 4.8 million.

28 ~’ NACA RM A52C10a

Figure Ii — Muxihwm. fift+hg ratio for friungulw wing wifh andwithout missile models stcwect externally us a function of IUoch

number. Reym/ok nwnber, 4.8milhn.

.

..

.

.

-..

NACA-hnglq - 6-26-5?, . S50