1 .

...--..CA

Copy No. 257

RM No. L6L24

RESEARCH MEMORANDUM

DRAG OF A WING-BODY CONFIGURATION CONSISTING OF A SWEPT-FORWARD

TAP ERED WING MOUNTED ON A BODY OF FINENESS RATIO 12

MEASURED DURING FREE FALL AT TRANSONIC SPEEDS

By

Jim Rogers Thompson and Charles W. Mathews

Langley Memorial Aeronautical Laboratory Langley Field, Va.

CLASSIFIED DOCUMENT CLASSIFICATIO~ C~A~GED TO

Thill document contains clusl!1ed informa.tio ,1 alfeeU .. "'" National Defense ( f th. UnI' I~CLASSIFIED Slates within the meaninC of the Espionage Act, ":SC 50:31 and 32, Ita trana:m1a&1on or t.h

DATE 8- 8- 54 revelalion of Ua contents in any maDDer Lt

unauthorized peraon fs prohibited by laW"'UT')'QUI o:O~~I:;.:n:l~ed ~~ .:~~ 11 n TY J. VI. .ervtces of the UnIted States, appropriate clvtllan officer. and employees of the Federal Government who have a leatUmate lateres)-. 'hereln, and to UnltodSta ••• C'tl2enooflalo ...... HANGE:#234 5 loyalty and dl.::retlon whO of beCeastty muat be lnformed tbf-reof.

CltO:YLEY

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

WASHINGTON

March 13, 1947

W. H L.

NACA RM No. L6L24 CONFIDENT~

NhTIONAL ADVISORY COMMITTEE FOR ABRONAUTICS

RESEARCH MEMORANDUM

- ---- --DRAG OF A WING-~ODY C01;FIGURATION CONSISTING OF A SWEPT -FORWARD

TAPERED WIN(} y..oUNTED ON A BODY OF FINENESS RATIO 12

Y.EASUT',sD millING FREE FJ LL AT TRANSONIC SPEEDS

By Jim Rogers Thomps on and Charles W. Mathews

SUMMARY

The drag of a configuration cQ:).sisting of a body of fineness r atio 12 vii th staDihzi::lg tail 8tU'faces and a 12 -percent ··thick, 300

swept -f 01'wc.cd vTing >aving a n aspect ratio of 4 and a t aper ra~:io of 2:1 has beer! measured at transonic specQs by the f:;:-oee-fall nethod. Tile t otal dra and t.."".J.e drag of tbe w-:i.ng were measured separately . These measurements) which were made as part of the NACA r esearch program to de terIl'j.ne optimum aerodynamic shapes and configurations fOT use in the t r ansonic and supersonic velo~ity r anges, show that the dl'ag of t J:le complete config1.U~ation r ose almost linearly from 0.07 of atmospl1eric pressure per uni t · of frontal area at a Mach n~1.D1ber of 0 · 90 to 0 · 30 of atmospheric preDsure at a Mach nVIDoer of 1.02. 'l'he d~ag of the wi ng r ose similar ly from 0 .047 of atnosr-heric pre s ure per l:1"1i t f r o;:}tal p,rea at a Mach number of 0. 91 to 0 . 30 at 0.98 3n d t i1e!1 i r:cre;1.sed mo:re slowly to 0 . 34 of atmos};':.er5.c l')ressure at a Bach n'uuber of 1.02 .

The presence of the swept -fOri-Tard wing resulted in a large unfavora ble i nterference effect on the d.rag of t1:1e body -tail combinat ion. These par ts experienced almost twice tho drn,g me::tsured i n previous tests of an identical body -tail combination wi thout wings.

INTRODUCTION

The ImcJ. .. i s te sti :c.g a s ories of wing -i.;ody conf igurations -;:)" the free -f all method (ref':)re::ces 1 and 2) t o investigate the t ran:'lOni c dra£ cha r R.cter istirs of pos sible E,lrplane al':;:-allgeL10:lts h<J:v~r.g diffe~~ent co:u.bina~ions of -..ring z '{CGll , t&'p,;r, and t:bickne8:.;. To l1l?\.~e

t~:e 8xperi " ental data available a.s s oon as pC'ssib13, tbe ::''3 6u1 ts of each test al'e being :p'J.blish fCJ d as soon as they have "!J0on 8 Tra l u8. tt;d.

CONFIDEITTIAL

2 CONFIDENTIAL NACA RM No. L6L24

'1'~le lY:'esent paper reports results obtained for O":le of t he sericsj a config~rctio~ consisting of a 300 swept-forward wing mountec. on a boo..y of fineness ratio 12 whose transonic drag charactcristiGs were known fro~ previous tests.

The resu.2..ts obtai ned from this test are presented as curve s showing the variation of erag coefficient wIth Macn number for the complete configuration and for its corr:ponent parts . The results are 8o:r::Ipa-r'ed wi "'vh those for an i d.entical body without wings reported in reference 3 and for r ectangular and swept-back airfoils mounted on a cyllnd:i.·:i. cal bod.y :reported in references 2 and 4.

APPARATUS AND METHOD

'J.~st . body .- Tha general a r r angement of the test configuration is shown by the photographs (figs. 1 fu~d 2), and its details and d imensions by the dr.~wing (fig . 3) . The body and tail "Tere identical wi t h the configuration of fineness ratio 12 whose tests were reported in refer ence 3. The wing had a sweepforward of 300

,

measured at the quarter-chord line, and NACA 65-012 sections perpen­dicular to t his line. '1'he t apar ratio was 2 : 1 and t he aspect ratio, based on the wing a:rea includir..g that submer ged vTitlli n the body, was 4 .0, The wing enter~d the bouy behind the maximu~ diameter

t hrough rec tangular slots 2t by 26~ inches and was attacned to a

spring 'bal ::mce iTi tb in the '0ody . C'hese slots vTere fille D. !-jY SIndll wanden olocks mounted on the \fing roots and shaped to preserve the

body concaur. Clear a..'1.ces of about '3\ i nch were provided so th?1. t the

end pla tes did not r ub against the side s of the slots a.s the wing balance deflected under drag load .

Me.ill:!}k1Z.ELment s_. - Measurement of the de s ired qua'"lti ties was accomplished as in the :previous tests (references 1 and 2) thr ough u s e of the N.'i.CA radio telemeterii g s.fstem and !'adar a:r.d photo ­theodolite ea.uipment . The following quan.ti ties wer'e record.ed at two e8parate ground stat lons by the t e13meter' i ng system:

1. The force exerted. on the body by the wing a s measu=sd by a spring balance

2 , The t ot a l retal'':a tion of the coml,lete configu::.'C\. ti :"1U as mr,asured by a se'"};:; :' ti ve acc31erometer alined. wi t~l the lo~git~dlnal axis of the b ody

CONFIDENTIAL

NACA RH No. L6L24 CONFIDENTIAL 3

3· The total pressure at an orifice l oca ted at the nose of the test body as measured by an aneroid cel l

A time history of the position of the body ,.,i th respect to gr ound axes durlng its fall was r ecorded by radar and phototheodolite equipment, and a sur-vey of atmospheric conditions applying t o the test was obtained from synchronized records of atmospheric pressure , temperature, and geometric altitude dt~ing the descent of the airplane from which the test body was dropped . l'he direction and yelocity of the horizontal component of the wind in the range of altitude for ,,,hich data are presented were obtained from radar and phototheodolite records of t he path of the ascenston of a free ba1100::1 .

~eductiop of data . - As in the previous tests, the veloei ty of t he body ",1 th respect to the ground, hereinafter referred to as ground velocity, during its fall "as obtained both by differentiation of the flight path determ:i.ned by radar and phototheodoli te equipment and by int egraMon of the vector surn8 of gravi tional acceleration and the directed retardation measured by the longitudinal acceler­ometer . The true airspeed \oTas obtained by vectorially adding the ground velocity and the horizontal "lind velocity measured at the appropriate altitude .

The total drag was obtained by multiplying the retardation ae (in g units) by the \om i ght of the configuration. The ... .,ing drag Dw was obtained from the relat ion

\"here

R measured reaction between wing and body, pounds

Ww weight of wing assembly supported on spring balances, pounds

The dr ag of the body -tail combination was obtained by subtracting the drag of t he wing from the total

The atmospheric pressure p, the temperature T, and the appropriate frontal a:;.'ea F were combined with s imultaneous values of true airspeed and drag to obtain .JL ratios for the cor:rplete

Fp configuraUon and its component parts and the Mach number M.

CONFIDENTIAL

4

area

CONFIDENTIAL NACA RM No. L6L24

Values of cOllvent:'Lonal drag coefficient based on frontal CDr were obtained from the relation

c~

D Fp

:---:x M2 2

where the r a tio of specific heats 7

coefficients based on plan ar ea CD was taken as 1 .4. Wing drag

wer e obtained by multipl ying

C~ by the r atio of wing f~ontal area (projected f r om the line of

maximum thicl;nes s ) to plan ar ea . Areas used did not i nclud.e those anclosed by the body .

Mach number wa s also obtained from the total-pressure meaSlITement by us e of the relation

M = I ~)~;l .1

\ I ------.. -._.-..

'Y - .L } -_._----, 2

where H is the measured total pressUl~e and the other symbols are as previously defined . This expr8s sion does not include a cor rection for the los s in total pressure through the normal shock wave which would appear in front of the orifice at supersonic speeds , as at t he low supersonic speed.s attained by thts te s t body t he correction Vlould be negligible.

. BESULTS AND DISCUSSION

A Ume history of important Cluantities obtained in the pr esent test is presented as figure 4.

The ground velocity ob taine~ for the t est body from the acceler­ometer data is shown on figure 4 as a (lashed line . The test body was tracked by the rad.ar and phototheodolite equi pmen t duri ng t he entire drop ; hm-rever J 6.ue to r el a tively poor vis ibilit y condi Mons and rough tra ck ng the theoc'coli t e photographs ; vrhich norma lly allow the data to be cor r ected for small tracki ng errors, were obtained only f or about 6 seconds near the Gnd of the drop. The ground

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-------------------~------ - -

NACA EM No. L6L24 CONFIDENTIAL 5

velocity computed from the radar and phototheodol ite data during this period is shown by the test points . The data from the two different sources agr ee and the accelerometer data. corrected to true airspeed by use of the wind data, were used to compute the Mach number. Both this Mach number and the Mach number determined from total­pressure measurements are plotted on the time his tory. They differ by a maximum of -:0.02, vThich is ,.,ithin the expected limit of accuracy of the pressure measurement . The Mach number determined from the true airspeed data was used in the remainder of this paper and is believed accurate within -to.Ol.

The r esults of t he tes ts are summarized on f igure 5 where curves are presented which shm., the variations obtained for ~- and C~

for the complete configuration, for the wing) and for the body and tail. The wing drag coeffiCient CD is also presented on this figure.

As the spring balance with which the io1ing drag force was measured must withstand the high drag forces occurring at supersonic Mach number s and high static pressures (low altitudes), it is neces sarily rela ti vely ins ens it:L ve to the small drags occurring at 8ubcritical Mach numbers at lov static pressures (high altitudes). The drag pe.rameters are ther efore less accurate at the lowest Mach numbers for which data are presented than at the highest speeds

attained. At M = 0 . 85 the ~ data of figure 5 are believed Fp

accurate within :0.008, -to.oo6, and -to.02 for the complete configu ­r ation, the ,.ring, and the body and tail, respectively . Corresponding values at M = 1 .02 are -to .004, -to .003, and ±0.01 . The CDF and CD values are somewhat less accurate due t o the introduction

of the Mach number values into the computation, t he uncertainty in CD for the wing being about to.OOl from M = 0.85 to M = 1.02.

In an effort to obtain the drag data as accurately as pOSSible, the maximum balance deflection was chosen to correspond closely to maximum drag estimated f r om available information . The actual ,.,ing drags obtained were consider ably larger than estimated and as a r esult full balance deflection occurred at M = 1.02 at an altitude of 12,600 f eet . (See fig . 4.) No significant data were lost, however) as the high drag of the test configuration prevented it from attaining a Mach number greater than 1.03.

Figure 5 shows that for the complete configuration the drag rose almost linearly f r om 0.07 of atmospher ic pressure per unit f r ontal

CONFIDENTIAL

6 CON]'IDENTIAL NACA RM No. L6L24

area at M ,= 0.9 :to 0 . 30 of atmospherlc pressure at M = 1 .02 . The .. , ing drag. ::;'ose s i'2:.ilarly froD: 0 .047 of atmospheric pressure per unit f::"ontal area ' at' M = 0.91 to O· 3C at M = 0 . 98 and t hen i~creE'.sed more slow'ly t,o 'O.34 , oi"atmospheric pressure at M= 1 .02 . The dr ag of the l),ody"and ta 1, 11hich ivas a.etermined by subtracting the vring drag from the total drag, · r.o~e lin~a!'ly from 0 .08 of atmoc.pheric pressurr3 at M;;:: 0.85 to ' o .16 at M = 0 .975, then inc'..'eaced abruptly ' 'to 0 .25 of atmosJ,.ber'ic p~essure at M = 1 .01. Cons::Lderable unsteadiness' was pressnt in the'd~ag valu0s at Mach nu:nners bet-w3en 0 .90 and 1 .00 . This uIlstet',i:.iness is beU8,,-ed due to disturbe d flow conditions a t the wing-body jlllcture .

''J:h -; d,j,'36 :v-esul ts obtained for the swept-forwar d wing and r esults of previous tests of rectangular plan for m 8...11d 450 sweit -'Jack plan-' for':'l1 aj.rf'oi'.s mountec. OIl C;) l1ndrical test bodies are cO:Jl:':'.3,rec. in figu:-e 6. ~~Fl, airfoil secti on normal to the qua.:.rtel' -chord lin"}, t he s-;..eop ane;le, the a t;pec t. ro. tio, and the reference from vihi ch the da t a ,.ere tal'.:en are given in the flgure . Tha drag l'ise measu'l:'ed in f.be

o ' .IU:'esent tests for the 30 swept ·forVle.rd, J2 -percent -thick airfoll of aspeet r at'io 4. and taper ratio 2 : 1 agrees cl osely -with t lle drag rise for the unsvlept 9 -percent-thlck airfoil of as:f)ect r ati o 5. 1 ( ~efere.nce 2) .

The r esults , of previous' tests presented in figur e 6 illus t r ate thfl effects. of thicK.."1ess <ind aspect ratio for rectcm.gular plan -form airfoils a:ld t~le effe::;t of ' h5° svleepback for an unta:ge1'ed airfoil . The se reuulta are not directly cO.1lpara-ole wlth the re tfi.J.lts of the p:;:-s.;ent tost, ~':bvreV'e.r, because t:~es e airfoils '-Tere unta1'0r ed, ,,;e':'e of d1f f"l r ,':::nt tbicli:ness .. and ,';rere !'lounted on long cylin:i, .-i cal . odies Vilich they ent ered throue,b open r 3ctangular slote . Consil't.:;ration of the flo-w about , the body 8110'\o:s that in the pres6P ,t test a J?ort ~·.on of the wing aear the root was in ,a region of increaSed vele ,: 1 ty dur·) to the curvature of the fil1eness ratio 12 body. The drag of t hIs por t ion of the wlng would ther efore begin to rise at a 10'!:Ter f r ee-s tre8.J1 Mach nu.n:.ber than. t'hedrag of an id.entical wing mounted on a long cylin.J.rical body where t.he exces s velocities ar e negligib le. For conr'igurat:;'ons 8,imllar to that tested, the drag 0-:: a tapered wing would te:-.d to rise mere abruptly than the dr ag of an illltaper e d wins bince a l al'ger pa ::.-t of the ,.ing area ,wnld be loca ted in the region of a?prsciable excess velocities.

'l"he drag of the bod.y-tail cOlJ1Ji.nation, obtained by sub t racting the drag of the vring from the ne<J.sured total drag, is COr:.p~l'ed. i::.l fi3t:''fe 7 w J, th the drag 0:: en i 'lentical 'body -tai l oillb:'n~. c,~ on te8teu. wi thout wir s s (refe:i:'ei1Ce 3). This comparison sho-vTs a 1c..rge 'U:l:t'avor­able interference effect; the d:cag of the body-tail cOID~')inatlO~1

CONFIDENTI Al,

-_._ - ---

NACA m No. L6I,21~ CONFIDENTIAL 7

being a lmost doub led due to the presence of the wing . Presumably this large, rather l'.:lstead;y interference drs.g res ults from dist.~:cbed flow conii tions a--c the wj.ng-fuselage Juncture.

CONCLUSIONS

The drag of a cOlU'iguration consisting of a body of fineness ratio l2 i ch stabiliZ:ill~ tail surfaces and a 300 sW'ept-forward wing has been IDflasur8d at b'8J.lSO!llC s peeClo by the free -fall method . Tho t.otal c}:~ag and the drag of '~he wi g wer'.; meas'LU'ed separately. The drag of tile cOLl,dlete coni'i.guration r ose almost ·l1nea:r. :~y fr Jm 0.07 of atmospl:er:.-c press~l.re per uni t fl'ontal area at a M..9.ch m:.!Jlber of 0.90 to () . 30 of atmosph0Yic pressure at. a Mach number of 1.02. The drag of t hp, w'ing ' r Ci:Jfl similarly from 0.047 of atmos pheric p:t'essure per unit frontal ar'3a at a Mach numbe_ of 0.91 to 0 . 30 at 0.~8 and then increased more slmvly to 0.34 of atmospheric pressure a t a Mach numoer of 1 .02 .

The pY8SenCe of the swept -forward wing resulted in a large unfavorable interference effe ct on the drag of the body-tail c om::> ina tionj these :;:)a~ts experI enced almost twice the drag measured. in l.revious testG of an identical boily-tail c Oillbiua.tion .... i thout wing3.

Langley Memorlal Aerona~tical La~ ~ratory Na tJ OJ.)al Advi!:;ory Committee f o::, Aeronautics

Langley Fiel('. , Va .

CONFIDEnTIAL

8 CONFIDENTIAL NACA RM No , L6L24

REFERENCES

10 Ba iley, F . J., Jr., Hathews, Charle s W. J and Thompson, Jim Rogers: Dra g J:4easurements a t 'llr a ns onic Speeds on a Freely Falling Body . NACA ACR No. L5E03 , 19 45 .

2 . Mathe",s ; Charl e s W. , and ':1:'hompson, Jim Rogers: Comparative Drag Measurements at Trans oni c Speeds of Rectan g1.1.1ar and Swe:pt-Back NACA 651--009 Airfoils M01mt ed on a Freely Fa lling Body.

NACA ACR No o L5G30 , 1945.

3. Thompson ) ':;' im Rogers, and Mathew's, Cha rle s W.: Total Dra g of a Body of Fine.Cless Ratio l ? an.d Its Stabiliz i ng Tail Surfac es Measur ed Dur i ng Free Fall at Trans onic Speeds. NACA CB No " L6D08, 1946 .

4 . Mathews, Charles W. , and Thompson, Jim Rogers : Drag Measurements a t Tranqonic Speeds of NACA 65-009 Airfoi l s Mounted on a Freely Fa lling Body t o Det e rmine t he Ef fects of SweepJ ack and Aspect Ratio . NACA RM No . 1EiKo8c , 19 4'7 .

5. Thomps on, Jim Rogers, and Mars olD.e r, Be rnard. W.: Com.parative Drag Measurements at Transonic Speeds of an NACA 65-006 Airfoil and a SY1J1.:J1e t r ical Circular-Arc Airfoil . NACA RM No. L6J30 , 1946.

CONFIDENTIAL

NACA RM No. L6L24 CONFIDENTIAL Fig. 1

Figure 1. - Three-quarter front view of test configuration.

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'-------- -- -

NACA RM No. L6L24 CONFIDENTIAL Fig. 2

•

Figure 2. - Top rear view of test configuration.

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

L ANGLEY MEMORIAL AERONAUTICAL LABORATORY - LANGLEY FIELD . VA .

CONFIDENTIAL

- --.-- - -----

144

65~ 120

6~~1 \ \ \ J

~ ROTATED 45° "- i -= _ /<)

t ~~ TAIL SECT ION

>-

~ ----~

W I NG SECTION (A-A)

01 MENSIONS, 5Q.Ff BODY COORDI NATE.S, IN. BODY FRONTAL AREA--- - 0 .545 W1N G PLAN AR EA - - - - - - 8 .000 WIN G FRONTAL AREA--- .767 TAIL PLAN AREA ------\.2. 32 TA I L FRO N TA L AREA---- .074 TOTAL FRONTAL AREA - - - - I . ~66

WING- SECTION COORDINATE.5, %C TAIL-SECTION COORDINATE.5, IN. (N AC A 65-012 SECTION) (NACA 16 SERIES AIRFOIL)

X Y X Y )( Y )( Y X Y

0 .000 0.000 20.000 4.975 65.000 4.38) 0.000 0 .000 1.350 .1 22 .~OO .923 25 .000 5.406 70,000 3.743 .0:; E .029 1. 8 00 .1 32 .750 1.10 9 3 0.000 5.716 75.000 3 .059 , I I J .04 I 2 .250 . 135

1,250 1.387 35.000 5.912 80.000 2.345 . 2. 25 ,056 2.700 .1 .J I 2 .500 1.6 75 40.000 5.997 S5.000 1.030 .336 .068 3 .1 50 .1 \8 5 .000 2..606 45.000 5,949 90.000 .947 .450 .or6 3 .600 094 7 .500 .3 .172 SO.ooo 5.7 '57 95:000 .356 .6 75 .09.3 4. 050 .0:; 7

10 .000 3 .647 5.5.000 5.412 100.000 .000 .900 . I 0.5 4 . 275 .032 /5 .000 4.402 60.000 4.943 1. 125 . I 1 !5 4 .500 . OO~

L , E. RADIU S 1.000 L , (. RADIUS 0008

X Y X Y 0 .000 0 .000 46.000 4 .87 6

.600 .2.77 54.000 4 .97 I

.900 .358 60.000 5 .000 I. ~OO .5 14 66.000 4 .955 :! .Ooo .866 72.000 4 .826 6 .0 0 0 1.446 78 .000 4 .610 9000 \ .936 84.000 4 .274

12.000 2 . ~65 90.00 0 3 .754 \8 .000 3 .11 2- 96.000 .' 3 03 I 24.000 3 .706 102.000 2 .2 22-,)0 .000 4 .1 56 108,000 1.3.50 36 .000 4 .489 114 .000 .5 26 42.000 4 .7 19 120 .000 000

NOSE RA DIU S 0.0 150

CONFIDENTIAL

NATIONAL ADVISORY COMMITTEE FOIl AERONAUTICS

Figure 3. - General arrangement and, dimensions of test configuration. All dimensions are in inches.

~ o ~

~ ~ . t'-i m t'-i I:\:l ~

f-xj ....... (J'q . w

Fig. 4

40O:J

3600

32(}O

280'0

<:::24O:J

~ ""-'2'0'0'0

600

II

10

NACA RM No. L6L24

CbNFIOENT!AL 40

:co::x Pxo --- 36

56v

26

'8u

44",

'Ou

8u

t::XA.

4u

cUV

o -8 -4

~

I>~ 0'? II ~

/ ~ 28 AN/Tude/ ~ /

~ 1/ 24

/ ~O ~ V

/

~ 16

/ /

V ~ [7 12

rota/ pressure ........ r-.......... / /

V V ~ 8

/ /"

V V ~ 4

V ~ --I-~ I--

~Sfaflc '0

f---~ - preswre -

---I-I..----" ~

Tempera/ure V "-.

~ ~

-----~

~ t:>- -------~

~ '"

- -..xtJ

Ground ""Ioc;fy (r07 radar and phololheodolde dolo--......." ~ , , ";; ""-

(j()')

AI(spee~~ --:;~ ~ 7tO

Ground velocdy from

~ V cP

acceleromeler dafa~ I---- "'" i---

~ ~ ~ ~ r-------.l'1ach number --.:: !W'O

/ from pressure dole.

~r> l1ach n(lmber frOI/l '0

~ /' £ Oirspeed- lemperafure ./ V 1- ~ /' ;/

dolo v 0'0

- -- ---- ./V V

300 Toiol drag

~ ~ k-HATIOH4L AOV/$OAY

'IinuRlt1WTICs

"~ ~ V ""-~Wlngl drag

o 4 8 2 ~ ~ ~ ~ 2 % ~ # ~ 2 Time offer release, sec

CONFI DEN TI AL

Figure 4. - Time history of free fall of test configuration.

NACA RM No. L6L24 Fig. 5

.~

CONFID~NTI~L q'-"'" \J~ 04 1:.. +-~ Q C s::: ~ 0 "-.l_

~Q IU c .02 (J (J \.)

Cj)] ~] . ~'-=

L r--('

/ I- Comp/efe confi5urahon -- Wing only _

/ ----- Body and tal I on Iy

....-/ v'

.6 0

r -....... I~ ..... -

-.j ,

/" f \ , v

fo--- .. " VI ...,-/ - /

-~--

.4-

/,," v

I;;' v

,... .. -....

I

~-I _rJ

. 2

-8 -;1 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

I L o .8 .9 1.0 /.I 1.2 .7

CONFIDENTIAL t1ach number

Figure 5. - Variation with Mach number of drag coefficients and D/F p

ratios for the complete configuration and its component parts. For each curve the drag parameter is based on the area of the specified component.

r----- --

l

--~ .OG ~

Q'+­'J () ;...<\ ~ c: ~ ~ C3 .04 \J c:-

'i::: \:) ~ -Ib ~ () c \..) ()

~~ .02 CJ V)

~ '-..:

o

CONFIDENTIAL

Section (NACA) 65-009 65-009 65-012 65-006 65-009 Aspect ratio 7.6 5./ 4 76 3.6

-- SweG'p 0 0 -30 0 4j -I?eference 4 2 Pre"8,t 'test 5 2

\ 1\ I I

\ \/ 7 I _

/ -1-1-----7

\ " \.- I -- / I

//~ -I.--

I

\/ l--- L ---- / i

/iJ V / / /1/ 1 -/ ---// - ---:

/~ ~ ..----.... -NATIONAL ADVISORY

COMMITTEE FOR AERONAUTICS I • I

.7 .8 .9 /0 II /.2 1.3 CONFIDENTIAL

Figure 6. - Comparison of wing drag results with those of previous tests.

~ ..... aq . CJ:l

~ () ::t>

~ ~ . ~ CJ:l ~ C\:> ~

II­~ "Cl c:: ~

(9~ 3 'J -E . .. ~ ........ C)

c: t'5 ~ ~ ~ ~ ~ - .2 'b-S! C) ~

\) ~

~§

~ ~ ./ C) ~

.f9

°7

.

I I I CONFIDENTIAL

lr'\.. __ r-- Present fest

J / I I

~ 1---1 \.1 I

,... I A --v '" / / ~ / . 1-/

/ '" ~

i'- Reference 3 f

I NATIONAL ADVISORY

CONFIDENTIAL C0j"ITTEj FOIl ArONAUTIICS-

I I I .8 .3 10 u 1.2 1.3

Mac h number Figure 7. - Comparison of drag results for the body-tail combination

with those of previous tests.

1.4

~ o ~

~ z o . ~ (J)

~ ~ ~

I-Ij t-'. aq . -:J