RESEARCH MEMORANDUM - UNT Digital Library/67531/metadc60199/m2/1/high_res… · to study tne effect of fuselage boundary layer on the internal-flow losses in tire i-LLet. The boundary-layer

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RESEARCH MEMORANDUM

LOW -SPEED LONG-ITUTDINAL STABILITY AND LATERAL-CONTROL

CHARACTERISTICS OF A MODEL OF A 400 WEFT-WING

FIGHTER-TYPE AiRPLhNE AT A REYNOLDS

NUMBER OF 9 X 10'

By Thomas V. Bollech and E. Neale Kelly

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NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

WASHINGTON February 3,1956

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NATI0W.L ADVISORY COMMITT€CE FOR MiBONAUTICS

Cm"SACTER1STICS OF A MODEL OF A 40° SWEPT-WING

By Thomas V. 3ollech znd H. Neele K e l l y

A n investigation w a s conducted i n t h e Langley 19-foot nressure tunnel on e. model of a 40° swept-wing f i g h t e r a i r p l m e t o determine modi- f icat ions which would eliminate the Ditch-up that occurred ?_ear maximum Lift din ing f l igh t t es t s of the airplane. Tce effects or" h igh-Ef t m-d stell-control devices, horizontal-tail locations, external stores, and various inlets OII the longitudinal chmacterist ics oI" t h e model were investigate&. For the most past , these tes ts were coducted at a Reynolds nuiber of 9 .O X lo6 and 8 Mach nunber o r 0.19.

The resul ts indicated khat from the s tEdpo in t of s ta 'bi l r ty the inlets should possess blunted side bodies. "he horizontal t a i l 1oca"led zt either the highest or lowest position fnvestiga-led improved the ste- b i l i t y or" the model. Three configurztions were found fo r the nodel equipped with the production tail which eliuirated the pitch-up through

' the l i f t r a r e up t o mex5m11~ lift and provided a s tab le stztic nargi-n- which did not vary nore than 15 percerlt of the mean aerodynamic chord through the l i f t range up t o 85 percent of n a x h u m l i f t . The three configurations are as follows: The production wir!!g-fuselage-tail combination w i t h an inlet similar to the production inlet but s d l e r i n plan form i n conjunction with either (1) a wing feme located &t 65 percent of the wing semisgarl o r (2) an Il.7-percent chord leadir-g-edge extension extending from 65.8 t o 95.8 percent of the wing sedspan and (3) the production wing-fuselage-tail cabbination with the production i n l e t e?d an 11.7-percent chord leadir?g-edge extensiofi extending from 70.8 t o 95.8 percent of the wing senispan.

2

INTRODUCTION

I" me S K L ~ L ~ I flcght t e s t s 01" a full-scale 40' swept-w:rg f ighter air- ;>:me reveeied thet the airplane possessed undesirzole pitchup chaxac- teris-kizs Eear maximum l i f t ( a t low as well as a t higk speeds). Ii vas believed that the undesirzble lcngitudinal s+,ability characteristics were zssocietej. vith the iocation of the horizontal te i l on t he a i rg lme m d t h e k r g e shoul-der-type inLe:s a t tke ving roct.

In order t G tietemir:e corrective modifications, a Yr-cfie; of %he air- p l a c e w e s tested ir- the Lmgley 19-fcot 9ressure tunnel. The model was explcyed to study the effects of: (1) changes in korizontal t a l l corfigu- ratior, (2) ckanges i n w1r.g root inlet configrat tons, , ar-d (3) niscellareous stall-control eevices on the lonql-klrdical s tzb i l i ty charac te r i s t lcs of the codel.

Ir- addition to the Troduction t a i l , whick hecl no dihedral and was located 28 percer-t of the semispan a3ove the wir,g-chord plar-e extended, three c l ternate t a i l conf ig~re t ions were investigated. One t a i l ccnfig- "ration reccmxzended by the Lzngley MDoratory was s. Zrooged %dl having -22" dihedrai and u t i i i z ing t he same point of a t t a c h e n t as the prodwtion I ta i l . O n -L1e basis or' &x malysis using the downwash dsta c? references 1 snd 2, it was believed that this configLire.tion trould mater ia l ly rehce 33" elinir,ate the high S i f t pitch-up. The other ei ternate -;ail configurations I were obtaixed by ettacking the produc%ion ayd the inverted drooped t a i l (22' dihedrzi) at the top of the ver5ical tai l .

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The effects of each of four ?airs of i n l e t s were investigated with the various horfaor-tal tail arzngements to deternine the eI"fect o r these conflgwation chaTges on the stabil i ty charac te r i s5cs sf the model. Or. the basis of these te&ts and from prockction considerations, an i n l e t which wzs sk i la r t o t h e production in le t bu t smaller i n Plan form was selected i n conjur-ction with the production t a i l t o be incorporated cn t'ne model t o r the Investigstion of stall-control devices on the longitudinal s tabcl i ty character is t ics of the xdeL. In ddl t ion , the e fzec ts of v a r i o ~ s wing devices on the longi tudina l s tab i l i ty chzac ter i s t ics of the nodel equipDed with the profixtion inlet and ta i l were also determined.

A brief investigation w a s m a d e t o dete-mil?e the la teral-control cha- acteristics of the m c d e l eqaip>ed with the Production fnlet and t a i l and also of the nrodel eqcisped w i % the production t a i l m-6 zn i n l e t s L x i l a r t o the proihct lon inlet blrt sraller in plan form.

TQe investigati.03 resorted herein was carried out f o r the most part E t a Reynolds number of 9.0 X IO6 an6 a Hach number of 0.19 thrmgh an mgle-05-attack range frox -4O to 30°. In an e f fo r t t o deternine the

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1

&rag coefficient, - Dreg O_&J

pitcMng-rroment coeff ic ient besed on a center of grevity

Cn

r a t e or" change cf pitchLng moment w i t h lift coeff ic ient

rate of cl?ar?ge of pitching moxeot with t z i l incidence

rolling-moment coefficient, corrected for model

aspmetry, Rolling mxent %SWb

y&wing-nment coefficient, corrected for no&el

zrngle of attack of w i n g chord plm2, deg

it tail incidence =@e in respect to the wing chord ?lane, deg

R Reynolds number based on the mean- Elerodynamic chord

¶o

sw projected w i n g area (excluding inlets), sq ft

free-stream dynan?ic pressure, lb/sq ft

wing span, ft

spanwise distance Eeasu-red from plme of symnetry, ft

vertical distance above chord plane extended along mean ae rodyndc chord, ft

inlet velocity ratio, - a

'0) exit total-pressure recovery

A inlet entrance area of both inlets, sq ft

H total pressure

P static pressure

Q volme rate of flow measured at fuselage exit,

v velocity, ft/sec

Subscripts :

i inlet

e exit

0 free strean

max maximum

z local

t

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cu ft/sec

MODEL

The codel o-f a 40° swept-wing fighter airplane Lnstalled il?_ the &gley 19-foot pressure t u m e l is sho7.a i n f i gu re 1. Fie model was of steel-reinforced wood const:mctioc an_d i ts principal dinensions and design Teetures are presented in f igure 2 and table I. A rigging diagraT of the nodel wing is presented i n f igure 3. The model W E S designed to allow t e s t s of high-l i f t End stall-control devices, horizontal t z i l arrangemects, external stores, and var ious inlets which varied in plzn i'orn!.

The pe r t inen t geme t r i c chzac te r i s t i c s of the inlets, devices, horizontal t a i l arrangements, and external stores -2 presented in f igures & t o 11 and tzbles I1 t o -Vi.

The h igh- l i f t and s ta l l -cont ro l devices collsisted of p l a in t r a i l i ng - edge flags, leading-eQe extensions, wing fences, end a leading-edge modification whfch increased the leading-edge radius end cmber or" the wing sections thus modified.

The trailing-edge flaps extended to 51 percent of the senispan and I had a chord of 22 percent of the w i r s chord measured g a r a l l e l t o t'ne

air stream. The flaps could be deflected 200 Etzld kOo pe-rpendicular t o the hinge l i n e ( f i g . 7).

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The leading-edge extensions were designed so that any desired span, chord, or spanwise location could be icvestigated d o c ! w i t h def lect iom of Oo and -loo measured i n a plane perpendicular t o t h e w i r ? ? 1eadill-g edge ( f ig . 6 and tables 11, V, z s d VI).

Details of the leading-edge nodification which increased the cam- ber and ieading-edge rad5us of the wing sectfons are shown i n figure 7. m-e various w i n g fences are shmn in f igure 6 and tables TI, V, and VI.

The various horizol?-tal tdl arrangements were c-rised of e i the r an undrooyed or drooped ta i l (-22O dikedral) a t tached to the ver t ical t a i l st 28 percelzt of the w i n g s d s p a n above the chord plane extended, and ul undxooped or Y-tail, (22O dihe&al) attache& t o the ver t ica l t a i l at 65 percect of the wing sexispan zbove the chord p1m-e extended. The droqed and Y - t a i l s had approxiua%ely 7 percent less projected area t h m the tails without any d i n e d r d ( f ig . 5 ) .

The aodel w a s equipped with p a r t i z l md f u - s p m ailerons which extended frm 51 t o 95.8 percent of the w i r g semisym and fron l3.k

equipped for a few tes t s wi th so l id and perforzted f lap-type sgoilers which extended from 13.4 t o 50 percent, of the w i q semispa and hzd an

t t o 95.8 percent of the wing semisgm-, respectively. The model was a l so

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f f average pro jectior? cf ? .E perceKt of the streanvise chord when derlec':cd - I 30a ( f i g . 8) . me area ZP the perforated spoiler was apprcxiaately

80 percent of the are& o r "she solid spoiler. Urless otheruise indicaLed all k t e r a l c o c t r s l t e s z s Tarere made with the allercns o r spoilers Eeflected on the le?: T ~ i c g .

The model was srovided with exhaust cones so tha t Yne inlet-exhaust area ratio could be vELried, thus providing a means by which the mss flow r a t i o a t the i n l e t s could be varied (fig. 9). The s t a b i l i t y dzta presented herein were obtained with the inlet exit full opeE. Flow survey rakes were installed a t t h e a p p r o x h t e engine cm>ressor face location md ir? t h e j e t e x i t for the purpose of measuring flow ra tes a t the abovz-mentioned locations (fig. 11).

Various boundary-layer diver ter plates were provided on the model t o study tne ef fec t of fuselage boundary layer on the internal-flow los ses i n tire i-LLet. The boundary-layer diver ter plates are shown i n figure 10.

Designation of T e s t Configrat ions

Listed below are the designations of the basic collrporrent par t s of t'ne model:

A wing-fuselage-vertical-tail combination

B external s tores ( f ig . 9 )

various inlets : (fig. 4)

DO production inlet

Dl inlet having e smeller plan form than DO wi-bh

D2 D l w i t h sidebody removed (sbula ted nace l le type)

leading edge swept back l 5 O

D3 serniflush inlet

DOS Do with spoiler on side body

Do w i t h incressed radius on s ide body

Do2 DO with approximate square side body

Y

V

i

t

t

*'

!

a

?z production t a i l - zero di'cedral t a i l locz-led a t 28 percent of the wing serr?isgan above the chord F l a e extemkd

droqed ta i l - shi lar t o the productios t a i l but having T A - .28 -22O dihedral loczted a t 28 percect of the w i n g senisgam

&ove chord plane extended

T - T-taiL - same as production t a i l but located at 65 percent .65 of' t'fie wi-% senispan above &or& p lme extended

T V Y-tail - similzs to the pro&uction tail but having 22O 65 dihedrd located a t 65 percent of the wir?g semispen

above the chord glane extended

High-lirt and s ta l l -control devLccs: ( f igs . 6 ami 7)

E leediw-edge extensior-s (fig . 6 )

I IeadFng-edge moaficzt ion ( f ig . 7)

F wing fences (fig. 6)

fif trail iag-edge f laps fieflecked (fig. 7 )

Detail designations of the component par ts are given i n figures 4 to 9. The nodel configurations descrFbed herein are formed by cmbinip-g the apgropriate model components with %he wing-fuselage-vertical-tail combination designated by t h e l e t t e r "A". For exan?ple, A +- T - f B represects a wing-fuselage-verticd-tail cabinat ion plus zero dihedral horizontal t a i l located at 28 percent of the w i n g semispan- above the chord plane extended plus external stores.

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TESTS AID COPmCTIONS

Tests

The t e s t s were conducted i n the Lan_gley 19-foot pressure tunnel wFth t'ne air conpressed i n t h e tu-nneel t o a pressure of apgroxhately 33 gounds per square inch, absolute. With the exception of the wir!!- fuselage-vertical-tail combination, the iavestigatior- w a s carried out at e. Reynolds number of 9.0 x lob and 2 Mach nmber of 0.19. IC the case of' the whg-fuselage-vertical-teil combinstion, force neasurenents were

J

obtained through a Reynolds nmber ranse fram 2.2 x lo* t o 11.0 x 10 . AlL t e s t s were conducted over an angle-of-attack range from -4O to 31°.

f 6

Longitudinel characteristics of the mode*l were d e t e d n e d f o r the xodel equipped w i t h and withcut various inlets, high-lif t and s t a l l - contrcl devices, horizontal t a i l arrmgenents, 2nd with Etnd without external stores. For the nost part , the longi tudinal s tzbi l i ty tests were cmdacted w i t h a horizantzl t a i l inciCiencc of -5O.

The la teral-control chzrecter is t ics were determined though aileron deflection range of tl8O by 3 O irxrernents for the outboard bilerons an^ 212O by 3O increments for the inboerc a i lerons. in the case of %he Zlap-type so l id and perforated spsilero, deflections of 4.7O, 9.4O, 1g0, 45O, 55O, m-d 90° were investigated. The a i l e r c c and spoiler deflections were maswed i n a plm:: per2eciiic-dzr t o the i r respective hinge lines.

Corrections

Corrections for wind-tuaiel jet-5antdary effects have been made t o t h e pitching, rolling, and yczwing moments. Corr4ctians for su2pori tare and interference have not been zpplied t o t i e data. However, these currections would not affect the cmparisons of “;..he h t a rxde herein. Jet-bcmdary corrections detedr-ed frop reference 3 and air-flow- nisalinenent correcticn of 0.lo, e s t k t e d on the bask of air-flow surveys and t e s t s of previous nodels, have been aFpiied t o the angle of attack md drag coefficient. The wag coefficients presented herein include the in-bernal drag of the i n l e t s .

Tables I1 t o V I swanarize the res-dts oj te ined frm the low-speed lorgi tudinal s tzbi l i ty tes ts . Figures 12 t o 34 present detail force and moment data of sone of ths more pertinefit results obtained durim the investigatioc of the lorgitudiinal stability an-d la teral-control c h r a c t e r i s t i c s of the m d e l . -411 of t he s t ab i l i t y data presented i n figures 12 t o 34 are for a t a i l incidence of a p p r o x h t e l y -50 u-xess otherwise noted. Tables V i 1 m-d V l I i present the individual ran- recovery p e s s u r e s that were determined a t the engine canpressor face location fo r inlets Dl a 6 D2 at several zngles of attack WKI, i n the case of i n l e t Dl, fo r severs1 boundary-layer diverter configura- t i o m . The variat ion of the mass-flow ra t io s and ran;-recovery characterist ics with angles of at-iack for the mr iaus ide ts a re p resente6 i n figxres 35 znd 36.

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NACA RM LjIcSL7

RESULTS APiD DISCUSSION

Lopgitudinal Stabil i ty Characterist ics

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EfTect of Reynolds number.- A few exploratory tests were conductea 00 the wing-fuselage-vertical-tail conbination t o deterrnine the e r fec ts of Reynolds m.zxiber. As h d i c a t e d i n figure 12, the e f fec t of var ia t ion i n Rep-olds nmber OE the pitching-moment clm,racteristics of the w i n ? - fuselage-vertical-tail combination from a Reynolds amber of 5.0 x 10 6 t o 11.0 x lob can, for all practical. purposes, be considered negligible. Although the effect of var ia t ion i n Reynolds number on the pitcbil?-g- moment charac te r i s t ics of the wing-fuselage-vertical-tsil combinatior- was found t o be snaU zbove a Reynolds number of 5 .O x 106, it did not appear conclusive that the same wo-dd be t rue fo r all test conlfigura- tions. Therefore, F t w a s decided t o conduct the investigation a t the highest test Reynolds mmfber poss5kle w i t h due consideration given t o economy of operation znd sustained operation of t e s t e q u i p z n t . Hence, the invest igat loa was conducted a t a Reynolds number 03 9.0 x 10 6 ra ther

r

thm- a t the h i g h e s t €?ey"olds nmber attainable of 11.0 x lo6.

Effect of inlets.- With the exception of varying the length of the internal duct l ines betieel? the leading edge or" the inlet aSa the leading edge of the w i n g , the in t e rna l ductirrg for the various iriLets was desi-gned t o allow ell of the verious illlets t o be i n s t a l l ed on the model without altering the in t e rna l duct l ines . It is assumed i n the following a s - cussion, therefore, that any variations which occur i n the longitudinal c h a a c t e r i s t i c s of the *.ode1 equipped with the di f fe ren t in le t s a re due en t i re ly t o the external effects of the i n l e t s .

In order t o show more clear-iy the e f fec ts of i n l e t s on t h e pitchtng- rnoaent c h a a c t e r i s t i c s of Yne nodel, figure 15 has been prepared, us ing the data of f igure 13, and presents the departure of the gitching-moment curve f rm t h e i n i t i e l l i r e s r i t y at Low l i f t tha t was obtair-ed for the modei x i t h and wit'nout t he i n l e t s . It wes discovered during the ir-itiel phases of the izvestigation t h a t the pitching-xanent characteristics obtained OE the model equipped w F t h the production hlet DO were not i n agreenent wi th those obtzir_ed during the investigation of the ' fu l l - scale d . r r , L a n e i n the Ares 40- by 8O-foot tunnel. It was recognized t h e t the prototype inlet incorporated on the riill-scale airplane differed from the prodGction i n l e t on t'ce model i n t h a t t'ne prototype inlet possessed a sharper side bokv than the well-rcunded side Sody or" the production Fnlet. Therefore, i n m ef for t to f ind an e s l a n a t i o n f o r t'ne cliscrepaxy ir- the two se t s oI" date, a s-ooiler was attached to the side

I ~ O C ? ~ cf i r l e t 20 Ln an attem-gt uc s h u l a t e . :o a reasonable extent, the aeroQnemic effect of SK inle'; possessing a skal-73 s ide body. !DE results ojtained :?it:? tke siml;la%ed sharp slde bod{ i n l e t DnS (rig. 13 ) were 2ocnd f,; ke i n sGYiciex5 sgreexent with the dats ob52ir.ed d-drLng ti:e Zull-scale Izvestrgatioc t o comlcde that tke differexes that es is ted ketween the t v o se t s of data obtained on the no&ei a??d the full-scale afrplme were z-ktributasie t o +,he difference ir the s ide bow shqes of t h e g r o t c t n e an3 t he p roduc t im h l e t s . It can be seen frm tke data gresentea in figure 15 Ynat t h e 2 3 d i t i o ~ of The sixclated snzr~ side bo6y i n l e t DoS resu lced i n a meximum destakii izing pitching momen-c of 0-173 whick was considerajly greater thm- that obtaired for t3e mcdel r;ithout inle$s. I n additior the a9gle-of-attack rexgc over which these increnents of destabilizLrg pitchil-g xoKclt existed Ioar i n l e t !los was ccnsiderably greater th&n for t'ne model wi th k i e t s 05:. It is evident from the fore- goillg discussion thzt a1 i n l e t having a sharp side body w m l d be de t r i - !-cental t o the lor@t=ldinel s t ab i l i t y ckxac te r i s t i c s cr' the airplane.

Exanination or' figure 15 reveals that, w i t h the exception of i n l e t D3, the addition of the inlets reduce6 t o sane extent the maximum increaent of destabil izing pitching moment of aFproxim3tely 0.111 that w a s obtained for the node1 without inlets a t an ansle or' attack of approxhately 21°. The greatest redilction, agproximately 0.030, i n the increment of destebilizing pitcbdng m0n;ent was obtained w i t h inlet D2. In the case of inlet D 3 (semiflush inlet) a slight increase i n t h e maximum increment of destabil izing pitching moment was obtained. In addition, it c m be seen that the increment of unstable pitching moment obtained for the model eq-aipped w i t h the various inlets and one feme progressively increased i n magnitude and extended over a progressively larger angle-of-attack range as the inlet s ize increased.

Presented i n f i g u r e 16 are the increEents of destabil izing pitching rnazent obtalned for the model equigped with various inlets and wing fences. Cmparison of the data presented i n figure 15 and f igure 16 indicates that a properly located fence generally reduced the aagnitude of the increments of destabil izing pitching moment by 75 percent for angles of attack below approximately 240. It will a lso be noted frm the data of fCigure 16 that the addition of one wing fence to the model equfpped with inlet D2, which has been previously shown t o p o v i d e s i p nlficant improvements i n the pitching-moment character is t ics , produced stable pitching-mcment increments throughout the -le-of-atteck rzngc above lgo. A t t q t s t o reduce further the magnitude and the extent of the iocrenents of unstable pitching mment that occurred for model equipped w i t h the larger inlets Do and D by using two w i n g fences

proved t o be somewhat successful as can be seen from the data of f ig - ure 16. However, even with two fences the pitching-ament characteristics of the model equipped with the la rger in le t s were s t i l l cot ES fav- orable as those obtained for the model equip?ed wi th i r Je t D2 and only one fence.

02

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J

Effect of horizolrtal ta i l locat ion.- Presented in Tigure 17 are the loP&itudi?ml chsrac te r i s t ics of the model equipped wikh various inlets

J and horizoctal ta i l arrangenents . The variations of dCmldCL w i t h l i f t ccefficieEt obtained for the various inlet ani horizontal ta i l arrangements are preseoted i n figure 18. Inspection of figure 18 indicates that o f the various horizortal tail arrmxenents investigated the Y-tail f r , regardless of the inlet configuration, was the only t a i l

arrangement which prodded cegative values of dC,/dCL through the l i f t range up t o or within 2 percent of i n the case of

inlet Do. However at or beyond the pitchirg-mmert character- i s t i c s becme ws tab le . In ell cases, the variatioc of dC,/dCL Filth liTt coefficient obtained with the Y-tail &id not exceed 15 percent of the m a n aerodynamic chord ug t o naxhun lift. The s d l e s t var ia t ion of dC dCL was obtained with i-rrlet De and was e q d to 0.0%

LL .65)

c4ax ch?.l.m chax

nl It cm- be seen from the data of f igure 18 that decreasing the ta i l

height by u t i l i z ing the drooped tail A did ~ o t e l h i n a t e the posi-

tive values of dC,/dcL that occurred near CT- w2th tne production

razze over wbich positive values of dC,/dCL occurred f o r the nDiiel

equipped with the Production tail so that i n the case of i n l e t s D2 and Dl - it is probzble that no pitch-up would be exgerienced in f l ight .

' .28 - tail. However, the drooped t a i l suf f ic ien t ly reduced the li-r"t.-coefficierk

A

Fx&rna%ioc of the r e l a t i v e a e r i t s or" the vaious horizolztal ta i l arranaenents tblrough a l i f t -coef f ic ien t rv lge ug t o 0.85 CT indi - cates that e i ther the T,:~ or the T .6; t a i l would prov5.de negative

values of dG/dcL for el1 inlet co-afig"s?tions except fo r F-nlet Do i n conjunction with t'ne drooped ta i l %-here positive values of dC,/&CL were obtained between E l i T t coeff ic ient of 0.8 and 0.86. The varia- t i o n of dC,/dCL that was obtained with the T A and T V tails

thlrough the usable lift range varied from 5 t o 20 percent ol" the Eean aero6ynanic chord depending 02 the inlet conr'iguration. The smallest var ia t ion of dC,/&C, through the usable lift range w i t h the drooped tail w a s ob-laiEed wi,t'n iKLet D, and was equzl t o O.OgC'. In the case of t h e Y - * a i l the smallest variation of %\&CL was obtained with i n l e t D3 m-d vas equal t o 0.06%

*ax

.28 .55

-

L 2

The va2Aes of dCln/Ciit obtailled a t zero angles of attack f o r tke var ieus horizontai-tall Izcatlcns c e l i s t ed In t h e foilcwinir; ta3 lc :

Horizontal-tail configuration

(a>

T - .28

I -0.0167

- .0187

aDe-Lermined f r a n dcte. G-F' P i g c e ;7(a).

Effect of various wing devices on the model equipped with the production t a i l and inlets Do or Dl.- The ef fec ts of various arrange- "- merits or coabinations of leading-edge extensions, w i n g fences an& leading-edge nodification OD the s tzk i l i ty charac te r i s t ics of the model equipped w i t h the production t a i l a& i-nlets Do and Dl were studied i n rn a%tenFt t o Find a wing configur&Aon which wodd provide stable pitching-mment chwacteristics through t'le l i f t -coeff ic ief i t rmge.

-

As an aid i n the selection of' the nost promising wiFg-device arrangement frm the stan&point of st8bili ty: a cr i te r ion hzs been adopted tha t the model m u s t not exhibit an adverse pitch-up tendency through the l i f t renge u~ t o C m d nust have a s t zb le s t z t i c margin wMch does not

vary more than 15 percent of the man .zerotiynamic chord tkrrough the iirt- c o e f f i c i e n t r a z e ur; t o 0.85 C It sho-dd be poil?-tzd out t h a t this c r i t e r ion was selected purely as a matter of convenience and sholdd not be construed t o mean Ynat this c r i t e r ion is a stm&rd s t a b i l i t y rcquire- rient. Also that the cor-clusions reached OIL the basis of thiis c r i te r ion mzy be somewhat d t e r e d if other cr i ter ia are uses .

h2X

ha=

. L

O f %he nany configurations investigated, severd configurations were founii which fulfilled the preceding requirements. These co-nfigura- t ions are: (1) 4- + Dl + T . 3 + 60 - F0.6~8, (2) A + Dl + T - + Eo.30(0.658 - 0.958), an-d (3) A + Do + T - + Eo.25(0.708 - 0.958).

The detail force data obtailzed with these coc3igwz;tions with and without flass deflected are presented i n figure 19. The variations of G / d C L with l i Z t coefficient, for these cor!!igurS;tions are presented i n figure 20.

- .28

.28

It is understood. that the production version of %he airplane i s t o be equipped wi th i n l e t DO, a leading-edge modiPication, m-d f l i g h t fences i n conjanctiozl w i t h the s t ra ight ta i l located et 28 percent OP the wizlg semispar? above the chord plane extended, wkereas t h e p s a s i t e vers io l or" the elrplane w i l l incorporate the &oog teil. I n l i gh t of t h i s understanding, it is or" i n t e re s t t o exmine the detaL1 force data obtained for the production and parasite versions of the airplane with f laps neutral and deflected ( f igs . 21 md 22). The var ia t ion of d&/dCL w i t h 1if-t; coefricient obtained f o r these configurations is presented in Figure 23. Figure 23 indicates that e pitch-up tecdency would exist neer

with f laps neutral as w e l l as f laps def lected for the production version. Droosing the horizontal ta91 22O redaced the positive valxes of d%-/dCL . new Cq- - but the redzction was not suff ic ient to e U -

inzte the pitch-up tendency. More s ignif icant thm the reduct ion in the pos i t ive vahes of dC, dCL the t w a s obtained with the droo-ped ta i l i s the loss i n s t a t i c mmgin that occurred. It w i l l be noted from the dzte t h a t drooping the horizontal tei l decrewed the static margin Tram approximately 10 t o 6.3 percent E with f laps neutral md from ZpproximEteU 10 t o 5 percent E. w i t h f laps deflected.

c4nax - a

-4

Effect of external s tores and f d e % mss-KLow ratios.- The effect of exbernel stores and inlet mass-flow r a t i o on the s t a b i l i t y of the model for vazious model configurations i s s h m i n figyes 24 and 25. It can be seen that the addition of excterral s tores had little effect on the l ineez i ty of the pitching-rimer-t cwves regardless of horizoctal tail locat ion or i n l e k codigurat ion. Hawever, it w i l l be noted %hat a slight decrease i n stztic margin was obtained i n every case that the external s tores were added.

Variations i n the inlet mass-flaw r a t l o appeared t o have no effect on t h e s t a b i l i t y of the model. The only significant effect of decreasing tne inlet mass-flaw r a t i o was a posi t ive trin shift.

c

14

Lateral-Control Cbracteristics P

Ailerons .- T'ne bts. gresented in f igures 26 an& 27 indicate that .c

the maxirnum values of r o l l h g mmtent obtEined with outboard ailerons was a,2_uroximtely 0.04 f o r a to ta l e i le ron def lec t ion of so f o r the model eq-dipFed with irrlet Dl znd for the model equipped wi th in lc t Do i n conjmction with the leading-e&ge codification md flight fences. In both cases, a 25-percent decrease in ro l l i l l g moment wes obtair.ed beyond a n -le of attack of 16O. Furthermore, i n the case of the xc&i equipped w i t h i n l e t Do Fn conjmction w i t h the leading-edge Eodification a-d f l i g h t fences, %he rolling-ncxent deta becaae very e r r a t i c i n m t u r e , znd i n soae instmces, zileron reversal occurred.

Compariscn of t i e r e su l t s of f i g w e 27 with those of f i g n e 50 Lndic8tes tb2-k no s ignif icant chawe in the ro l l ing mment was obtained by replaci-ng the leading-edge nodification an6 f l ight fences w i t h an ll .7->ercent chord leading-edge extension wMch extended frm 70.8 t o 95.8 percent of the w i n g smispm, (with flaps derlected in the latter case). However, when the outboard ena of the extension w a s noved inboard t o 0.858b/2 (f ig . 31) a slight decrease in C was obtained

and tine var ia t ion or' ro l l ing moment w i t h a zbove a- male of attsck of 16' became l e s s e r r a t i c w i t h l i t t l e or EO eileron reversal . Alihough no data were obtained, it is reasonalble t o e q e c t thzt an improvemnt in the var ia t ion of ro l l ing mment w i t h a would also 3e obtained wi th f laps neutral i f the shcrtened span of leading-edge extension was L.

enployed .

2,iaX

*-

inle cate 2 4 O

r B Ihe la teral-control k t a obtained on the model equipged w i t h t DO, leading-edge Kodificztion and f l i g h t fences (fig. 28) indi-

tot& def lect ion of the full-span- ailerons as was obtained with 3 6 O tht the sane degree of rol l ips effect iveness w a s obtained w i t h

to ta l def lec t ion of the oiltboard ailerons. As i?2 the case of outboard ailerons, t'ne vzriation of ro l l ing nonent with a for the full-span ailerons =hove a = 16O was erre.tic and i n sme instances aileron reversal was obtaimd. Therefore, as d g h t be expected frm tce data obtained w i t h Pall-s~an ailerans, it w i l l be acted frm a coEpLrison of the data preserted in figures 27 and 23 tht tne w e of differen-Li8,ll.y operated flags i n conjunction w i t h oztboard ailerons as EL lateral-control device appears t o offer sme &vantage over outboasd a i le rons a lme f ro2 tae stan53oTnt of rolling effectiveness.

Spoilers. - The lateral-ccctral character is t ics of 0.3b/2 spvl so l id and serforated flap-type spoilers are presented i n figures 32 and 33 f o r the nodel eq-digped w i t k i n l e t DG, leading-e6ge xodTficatior, and flight; fences. Camparison 03 the data yresented i n figares 32 and 35 reveals L %hat at low angles of attmk the rolling rrianent produced by e i ther so l id

$

or perforated spoilers qeflected 5fj0 was nearly equal-?o 50 percent"of the @lit-z noncE$Arodwed by an outboard f l a p t y p e a i l e r o c fo r a to t& aileron deflection of LBO. At high angles of attack both spoilers becane ineffective. The va&tions of C 2 with spoiler deflection at various

"" -

vlgles of a t tack 2;re preselzted i n figure 34.

Thus it czn be seen t k t spoi lers were inferior to flap-type ailerons from the standpoint 03 ro l l i ng ?ament produced. It 5s probable that s m - what better spoiler efzectivexiess would be obtained with a more optiauw spoiler arrasement.

The yzwing-nment data obtained with flap-type ailerons and spoi lers a r e i n accordaace wi th c m o n experrenee i n that .the yawirg n m n t pro- . duced by ailerons is generally unfavorable while that obtained with - spoi lers i s Tavorable over nost of the angle-of-attsck range.

- mterna l Flow Measureaents

Effect of b o m k y - l a y e r d i v e r t e m . - F i v e s 35 snd 36 snd tables VI1 a d VI11 present tjle i n t e r z " flow aeasiirenieEts obtained on the model

3xzaimtion of the dzta presented i n figure 36 and tables V i 1 and VIII indicates tbAt replacing the origiv21 boundary-layer diverter block with s p l i t t e r p l a t e s s l i g h t l y improved the inlet air-i'low charac t e r i s t i c s . The grea tes t impovement was real ized w i t h the smaller of the two spl i t ter p lakes icvest igated. The inprovenent that was obtained resul ted from a decrease ilz the localized losses w3ich occurred s.t the inner corners of the inlets.

CONCLUSIONS

AI? investigation hes been conducted i n t h e Langley 19-foot pressare

tunnel at a Rep-olcis mmber of 9 .O X LO on e model of a 40' swept-wing fighter airplane to determine modiXcatfom that would ixprcve the LOT- sFeed long ikd ina l s t zb i l i t y chwac te r i s t i c s of the z i rp lme. Tke hteral-control character is t ics of the model were elso determined.

6

L Prom the resul ts of the investigatiozt, t h e I"ol1owing conclusions

a re made: Q

1. The addition of an inlet w i t h a sharp side body imreased the destabil izing pitching nomemt tht occurred near Cq- fo r the model wit'roat irLLets, whereas a reduction in the destzbil izing pitchi-ng Eoment was obtained with i n l e t s having blunted side bodies. In addition the angle-of-attack range over which the increments of destabil izing pitching nonent exis ted for the model equipped w i t h a sharp side body inlet was considerably greater than for the model without i n l e t s .

-

2. The horizontal ta i l located st either the highest or lowest position investigated during the present tests improved the s t a b i l i t y of the model. The greatest improvement in s t sb i l i ty assoc ia ted w i t h horizontal t a i l aodification w a s obtained w i t h a "Y" t a i l (22O dihedral) located a t 65 percent of the w i n g semispan above the chord plane extended. This t a i l arrangemnt provided a s t ab le s t a t i c margin which did not vary nore than 15 percent of the I;?ea.n aeroiiynmic chord up t o maximum l l f t or within 2 percent of naximum lift regardless of tlhe inlet configuration. The ckooped tat1 decreased the range of l i f t coefficient over which the pitch-up occurred t o such an extent that it is probable that no pitch-up tendency w o u l d be experienced i n f l i g h t .

3 . 07" all tae arrangements of wing devices irwestigated on the model equippeO w i t h the p rohc t ion t a i l i n conjunction w i t h the proauc- t i on inlet or an inlet similar t o t h e production in le t bu t smaller i n p l m form, three were founci which eliminated the pitch-up and provided a s t a b l e s t a t i c margin w3ich did not vary raore than 15 percent of the L

nean aerodynanic chord up t o 85 percect of m z x i m m l i f t . The three configurations are as follows: The prodmtion wing-fuselage-tail conbi- m t i o n with inlet similar to the productioc i l l let but smaller in @an fom, Dl, in corrj-mction w i t h e i ther , (1) one w i n g fence located at 65 percenl; of the wing semispan or, (2) an 11.7-percent chord leadine- edge extension extending from 65.8 t o 95.8 percent of the wing senzfspan, and ( 3 ) the production wing-fuselage-tail cmbination w i t h the produc- t i on inlet and an 11.7-percent chord leadillg-edge extension extending from 70.8 t o 95.8 percelzt of the w i n g semispm.

c,

4. The s t a 5 i l i t y of the model was not zffected Eppreciably by the addition of either external stores or a change i n inl-et velocity ratio.

5. Beyond an angle of attack of 160 which corresponds t o approximately 80 percent of maxinun lift, a 25-percent decrease i n r o l L i w rroment was obtained for a l l Tlap-type silerons investigated acd i n the case of the =ode1 equipped w i t h the production inlet the r o l l i c g moment becane very errat ic i n nature and i n sone instmces a t leron reversal was obtsined. The addition of an 11.7-percent-chord leadfng-edge extension extending from 70.8 t o 85.8 percent of the wing semispar, resulted i n e ro i l ing mcments which were l e s s e r r a t i c w i t h angle of attack w i t h l i t t l e or no aileron reversal .

0

a

6. The ro l l ing monlent produced by E 50-perce3t-sexispm- solid o r gerzorated fkp-type spcile-r dei'lected 55O w a s ne&rl.y eqJal to 50 percent of' tile ro l l fng xomenc prodrrced a t low lift by an out i ioud f l a p tm.y-pe ai leron for a t o t a l aileron deflection of 16O. Beyond an mgle of e t h c k of lTG, however, both types cf spoilers vere ineffective.

*

Lagley Aeronautical Ukoratory; National Advisory Cormit tee for AeronaJtics,

Lmgley Field, V a . , Februvy 1, 1954.

1. Foster, Gereld V., and Griner, Roland F.: 4 Study of Several Factors Afi'ecting the Stability Contributed by a Eorizontal T a i l a t Various Vertical Positions on Sweptback-Wing A i r p l a n e Model. NACA FM LgHlg, 1949.

2. Salmi, Reino J. : Eorizontel-Tail Effectiveness and Downwash Surveys - for Two 47.7O Sweptbeck Wing-Fuselage CmTDinations With Aspect Ratios of 5.1 m-5 6 .O at a Reynolds Number or" 6 .O X LO . NACA 6 RE.I ~ 5 0 ~ 0 6 , 1951.

1

3. Sivells, James C., and Salmi, Rechel $5.: Jet-Boundasy Corrections fo r Complete and Semispen Swept Wirgs i n Closed Circular Wind Tmnels. NACA TN 2454, 1951.

TABLE I

GEOMETRIC CIWIACDISTICS OF THF MOEXL

A. Wing Assembly

1. Basic data: Root airfoil ( theoretical) , measured normal t o

Tip airfoil (theoretical), measured normal t o 0 .25 -chord l ine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.25-chord l ine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Angle of incidence, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . Geometric twist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sweep of quarter-chord line (true), deg . . . . . . . . . . . . . . . . . . . . Taper ra t io . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aspect ra t io (excluding in le t area) . . . . . . . . . . . . . . . . . . . . . . Airfoil thickness (parallel t o airplane center line,

percent c ) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sweep of leading edge (true) , deg . . . . . . . . . . . . . . . . . . . . . . . Sweep of leading edge (projected), deg . . . . . . . . . . . . . . . . . . . . Cathedral, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Root chord (theoretical), parallel t o a i r stream . . . . . . . . . . . . . . . Tip chord (theoretical), parallel t o a i r stream . . . . . . . . . . . . . . . . Mean aexodynm5.c chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . Location of mean aerodynamic chord, spmwise (projected) . . . . . . . . . . . span (projected) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Span (true) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . wing area (excluding in le t area), sq ft . . . . . . . . . . . . . . . . . . . . Area of w i n g blanketed by fuselage, sq ft . . . . . . . . . . . . . . . . . . .

2. D.imensions :

3. Areas:

NACA 6hn010

NACA &A010 . . . 1.50 . . . 0 . . . 40.00 . . , 0.378 . . . 3.45

. . . 8.10 . . . 4.2.51

. . . 42.56 . . . 3.50 44.577 in. 23.800 in. 36 3 5 in. 27.126 in .

~ 0 . 6 7 h in. * 120.900 in. 2

29.230 ~1 r . . .

, . . 4.554 5 4

r I

t I

TABU I. - Continued

GE0MR:TRIC CWCnISTICS OF TfE MODEL

3. Horizontal T a i l Assembly

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1. Basic data: Root air:Coil, measured n o m 1 t o leading edge NACA 61.~~009 Tip airfoi l , measured n o d t o leading edge NACA 64AO09 Angle of incidence . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Variablc

Sweepback (leading edge), deg . . . . . . . . . . . . . . . . . . . . . . . . . ko . 00 Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.00

Dihedral, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.59

Chord (constant) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14..hOO in. Mean aerodynamic chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14-.].cOO in. Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51.000 in. Distance from 0.25E of w i n g t o 0.256 of horizontal t a i l . . . . . . . . . . . . 69.356 in.

2. Dimensions :

3 . heas: Total horizontal t a i l area, sq f t . . . . . . . . . . . . . . . . . . . . . . . 5.022

!CABLE I. - Continued

GEOMETRIC CIIARACTEEISTICS OF TIE MODEL

C. VerticaJ. Tail Assembly

1. Basic data: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

I . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Airfoil, measured normal t o 0.25-chord l i ne NACA 64AOll Sweepback of c/b line, deg 41.27 Aspect r a t io 1.68 Taper ra t io . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.402

Root chord (theoretical.) . . . . . . . . . . . . . . . . . . . . . . . . . . . 28.739 in. Tip chord (theoretical) . . . . . . . . . . . . . . . . . . . . . . . . . . . . lO.gO0 in .

Vertical t a i l area, sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.861

2. Dimensions :

3. Areas:

D. Fuselage

Location of s ta t ion o (measured frm nose of airplane), in. . . . . . . . . . . . 14.805 Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 153.120 in. Maximum width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15.012 in . Maximm height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20.772 in. Frontal area, sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.749 g Fineness ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.59

Side area (excluding vertical ta i l ) , sq f.1; . . . . . . . . . . . . . . . . . . . 18.558 r

iz Volume, cu ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.499 i2

y- ? “I

r . I Y

-

c d c a

I P A B L E I. - Continued

GI’METRIC CIIARACTERTSTICS 017 TIB MODEL

E. Inboard Flaps

1. Bnsic data: ‘ ~ e . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Anglllar travel, measured i n a plane normal t o

Location of inboard edge, measured normal t o

Location of outboard edge, messured nortml. t o

Wing chord z;t inboard edge, measured parallel t o

Wing chord at outboard edge, measured paral le l t o

Location of hinge center line, measured n o d

hinge l ine, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . t‘uoel.age cenLer l ine . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1’LLselage center line . . . . . . . . . . . . . . . . . . . . . . . . . . . . fuselage center line . . . . . . . . . . . . . . . . . . . . . . . . . . . . fuselage center line . . . . . . . . . . . . . . . . . . . . . . . . . . . . t o 0.23-chord l ine . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Root chord, measured paral le l ‘GO fuselage center line . . . . . . . . . . . . Tip chord, measured parallel t o fuselage center line . . . . . . . . . . . . . Area of one flap, sq f t . . . . . . . . . . . . . . . . . . . . . . . . . . .

2. Dimensions :

3. Area:

Plain t r a i l i ng edge

. 0 to 40

. . 9.36 in.

. . 31.1J.i in.

, 11.1.65 i n .

I . 0 . 7 5 ~

. . 1.36

TAJ3W, I. - CmLinued

GEOMETRIC CIiARAC'JXRISTICS OF T€E MODEL

F. Ailerons

1. Outboard ailerons : (a) Basic data:

m e . . . . . . . . . . . . . . . . . . . . . . . . . . . Angdar travel, measured i n a plane normal. t o hinge

l ine, deg . . . . . . . . . . . . . . . . . . . . . . . Location of inboard edge, measured normal t o fuselage

center line . . . . . . . . . . . . . . . . . . . . . . Location of outboard edge, measured normal t o fuselage

center l lne . . . . . . . . . . . . . . . . . . . . . . Wing chord a t inboard edge, measured paral le l to fuselage

center line . . . . . . . . . . . . . . . . . . . . . . Wing chord a t outboard edge, measured parallel to fuselage

center line . . . . . . . . . . . . . . . . . . . . . . Location of hinge center line, measured normal. t o

0.25-chord l ine . . . . . . . . . . . . . . . . . . . . (b ) Dimensions :

Root chord, measured parallel t o fuselage center l ine . . Tip chord, measured parallel to fuselage center line . . Area of one aileron, sq ft . . . . . . . . . . . . . . . . (c) Area:

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

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

. . . . . . . . . .

26.71 in.

0 75c

P c *, .

TABU I. - Continued

GEmmIC CIIARACTERISTICS OF THE MODEL

F. Ailerons (Cont .)- 2. Pull.-span ailerons :

(EL) Basic data Type . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Plain Flap Angular travel, mcasured i n a plane normal t o hinge

l ine, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -18 t o 10 Localxi.on of inboard edge, measured normal t o fuselage

center line, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.36 Location of outboard edge, measured normal *to fuselage

ccntex l ine, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57.89 Wing chord at inboard edge, measured p ~ l r a l l c l t o fuselage

center line, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41.65 Wing chord at outboard edge, measured parallel t o fuselage

center line, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26.71 Location of hinge center line, measured normal t o

0.25-chord line . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0 . 7 5 ~

Root chord, measured parallel t o fuselage center line, in. . . . . . . . . . . . . . 9.29 Tip chord, measured parallel t o fuselage center line, in, . . . . . . . . . . . . . 5 .&(

Area af one aileron, sq ft 2.60

(b) Dimensions :

(c) Area:

"

T l L E I. - Continucd

GEOMETRIC CImCTERISTICS OF THE MODEL

P. Ailerons (Conc. )

3. Diboard. spoilers : (a) Basic data

Ty-pe. . . . . . . . . . . . . . . . . . . . . . . . . . . . Angular travel, measured i n a plane normal t o hinge

line, deg . . . . . . . . . . . . . . . . . . . . . . . . Location of inboard edge, measured normal t o fuselage

cen te r l ine , in . . . . . . . . . . . . . . . . . . . . . . Location of outbowd edge, measured normal t o fuselage

W i n g chord at inboard edge, measured p a r d l e l t o fuselage c e n t c r l i n e , i n . . . . . . . . . . . . . . . . . . . . . .

W i n g chord a t outboard edge, measured parallel to fuselage c e n t e r l i n e , i n . . . . . . . . . . . . . . . . . . . . . .

Location of hinge center line, measured parallel t o fuselage center line . . . . . . . . . . . . . . . . . . . . . . .

13oot chord, measured paral le l to fuselage center line, in. . Tip chord, measured parallel t o fuselage center line, in. . Area of one spoiler, sq f t . . . . . . . . . . . . . . . . .

c e n t e r l i n e , i n . . . . . . . . . . . . . . . . . . . . . .

(b ) Dhensions

(c) Area

4. Perforated Inboard Spoilers

should be as follows:

Area of one spoiler, sq f t . . . . . . . . . . . . . . . . . Area removed by perforation, sq f t . . . . . . . . . . . . .

This section is exactly the same as 3 except for 3 ( c ) which

(c) Areas

. . . . . . . . . . . Flap

. . . . . . . . . . . 0 t o go

. . . . . . . . . . . 11.64

. . . . . . . . . . . 40.94

. . . . . . . . . . . 34.97

. . . . . . . . . . . O.'~OC

. . . . . . . . . . . 3.23 . . . . . . . . . . . 2.75

. . . . . . . . . . . 0.37

. . . . . . . . . . . 0.37 . . . . . . . . . . . 0.07 m -1 i-

f c I. 4 N"

TABU I.- Concluded

G. External Tanks (450-gallon capacity)

Length, in . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75.47 DiamzLer, i n . . 8.81 Frontal area, sq ft 0.42 Angle of incidence, relative to fuselage center line, de8 -4.25 Spanwise location, measured normal t o fuselage center Ibne, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.18

Vertical location of nose of tank, measured normal -to fuselage center line, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -16.69

LongiLudinaL location of nose of tank, measured pa ra l l e l t o fuselage center line, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31-.25

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

I . . . . . . . . . . . . . . .

H. 'Pylons

Leading-edge sweep, re la t ive t o a l i ne normal t o fuselage center line, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30.0

Trailing-edge sweep, r e l a t ive t o a l i ne normal t o fuselage c e n t e r I l n e , d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30.0

Chord, measured along line -2' f r m fuselage center line, in. . . . . . . . . . . . . . 27.04 Thickness ra t io , measured along l i ne -2' from fuselage center line, percent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.25

Spanwise location, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13 -18

26

- I

- I Y

Ycrml leadlnpsdqa Chord-oxtsnslm

?ad:ua

I_

28 eaawmm@ NACA RM LjkB17

T L.S. d¶w:co

bu a a t

4s n

.p7 23.0

.o? 25.0

.oi 2lb.3

~ 1 8 22.2

.15 22.0

.4 4 . 0

.IC 25.0

- 2 ~ / b = 0.850

c

0.70tb/2

O.E56b/2 t o

e 2 Y b = 0.452

a

- Lending-edge

modifiortlon

edge radius 2 x normal leadlng-

Fencs conflgurltion

2y/b = 0.708

2g/b = 0.850

zy/b 0.402 - 2y/b 0.482 - 2y/b = 0.6p

Cm curve

CL

cm -.l

- .2 OF2 T I

%Aighaat angla o f t e s t

NACA RM LjhBl.7

". F",

I /

I - I - I

.I

.

T cI curve

31.c

* p r

32

Ii

.

c

33

34

- E/t = O.LC Z y f i = 0.i:t

c

35

z.

L

7

I

I

.

c

'

.

37

i

c

.

40

c

Y

TASU V.- Cor.tinued

NACA €04 L54BL7

Y

r

0.117,

0.1171

i

43

c

44

6 T 5.'. ea.1-a

1.059:

0.117~

45

r

0 .4 .8 1.2 CL

-.P 1

z? .O

22 .l

31.0

-.

23.1 I" -

2y/t = o.ge2 .

31.0

- 25.2

- 2y/t = 0.4Q - 2yf@ = 0.sp 1 "_ " - 2rfi = 0.708

2

L = g x 1G.q . T '.". .le I., -

;:.a:. - 2,/% = c.rjr

- 2,R = 3.7a

47

48

r T c - - spsn

0.65ebh

0.95eb/i to

0.652b/Z t o

0.958b/i

0.6 5!b/2 to

0.95Sb/2

T

- 2y/b J.LE2

23/b = 0.652

- 23/b = 0.6%

c

c

49

0.065

.203 *137

:% .495

C l v e r t e r Elock

p r i f i c e !IL - ?o ? - P .lunbe= 0

9, 9, I

:s I I 0.266 "GL3 -.316

105

T t

a = 10.6O

0.080 .141 .la

.Om

:taa

T Ssl l t ter Plnte No. 2

xi - Po 9,

0.959 ,393 .991 -591 .93

-.274 .7?4 .997 .995 .650 .ego ,931 ,995 -651 .a26

. 8 3

0.063 .122 *153

.a53 ~ 2 6 0 .1j0

F Dlvwter Block

0.1 b .It7 .127

:zt 222 .1& .162 .163

A39 .23l -325

.263

.377 .5co

.261 :@

:I22 -456

@.as -.071 -.318

0.2 1

-.3w - . o h

.

Spllttar Plate No. 2

pz - Po 90

0.063 .r22 .153

.a51

: J Z

- .010 “083 -128

” -

1 liivorter Block

Hi - =o

0.167 .160 .1 3 .178 .192 .205 .231

*13k

.181

.220

.266

-0.0 2 .Oil .241

- .071 .119 -261 1

Spli t ter Plate No. 1

Hi - Fo

.393

2 3 9 *lG

:kt; 573

.TO6

.1;47

. =23

::&

Spli t ter Plate No. 2

tk - yo 9,

-135 -:;g

-.083

.272

.112

7

1

i 2 z 9 10

100 101 102

a = o0

PI - Po 90

c = 10.6~ I c = 20.90

P I - 3, aQ

I Orifice Number

106

lox 10

109 110 111

Ill - Po 9,

o .290 -099 ,426

.123

519 "-

a = 10.6'

Hi - Po I

o .244

8499 ,101

.119

= 4 0 -"

1 Hi - Po

qo P t - Po 90

0.163

* O k 6 .1 9

- . lo1 "- -2 59

c

. "

55

57.797 ./4 ling -. /,

I I

. I L

Figure 3 .- Wing rigging diagram OS the mod.el. hl.1 dimensions are i n inches.

. "

r - 4

r+!

I

designation Tnlet

I

V i e r F-P

Figure 4 .- Details of i n l e t p1m. forsrs and contours. A l l dimensions are in inches.

I * .

! / 7 I

L,

.28 and

DaalgnaLlon: T . 3

"laWI A-A

Designation: T.3

Figure 5.- Details of Lhe various horizontal Lail arrangcments. Al.1 dimensions are in inches.

1- ,n 1

i Designation: T,g

/ """ "---+-""- + -- ""#-+."&-

view D-D

"

I W " t

.

(b) T . 5 and T . 6y

Figure 5 .- Concluded.

r *

"

Spanw! a8 Chordwise sxtunt location

(fraction, b/2) (fraction, c )

Section A-A (anlarged) typical

Symbol

1F

Leading-edge sxtsnrion configration

0 i . 1 7 ~

Symbol

E

1 --- - -- Inboard f l lght fenoa ay/b = 0.h82

Figure 6 . - Details of high-lift and stall-control devices.

”-

“ W e

Leading-edge modification

Section C-C (enlarged) Symbol

typical

Flap hlngo line 0.779~ - --,. ”railing-edge flap conrlguration

Designation: 6, = 20’ and/or 40’

\

Figure 7.- Details o f leading-edge madillzation end trailing-edge flaps.

L

I I

1

\-Alloron hings l ine 0.7790

Aileron oonriguration Spoiler oonfi.garation

Section A-A (enlarged) typioa1

Seotinr D-D (anlargod) typiorl

Figure 8.- Details of the various are

1ateraLcontrol devices. A l l dimsnsions i n inches.

w cn

Externa l s tore conr iwat ion Designation: 1.e. B \ \\

! Interchmgoabls

xhaust cone

E

Section B-B Saction A-A z I?

bJ

(enlargod)

% Figure 9 .- Details of external sbore and exhaust cone installation. A l l -s

dimensions are in inches.

\

.

"

Figure 10.- Details of i n l e t boundmy-layer diverters. A l l ahensions sre in inches.

I -. "_

70.920 -

-- - - orlfices (0.055 diam.)

Wall static

View B-B

0 5 Inches

View A-A Scale for views A and B

tube 8

L

.8

.6

-.4 .

.04

0

-.04 cm

-. 08

(a) C, and C, against a.

F i w e 12.- EZfect of Reynolds nu-ber on the wing-fuselage-vertical- ta i l conbination.

.6

.5

.4

.3

.2

./

0

c,

.04

0 Ch7 704

-. 08

-./2 -.4 :2 0 .2 .4 .6 .8 LO

c, (b) CD and C, against CL.

Figure 12.- Concluded.

. ./6

./2

.08

.04

. 0 -./2

- 0 -./6 Gn

0 -20

00

-. 08

-. /2

-.I6

-. 20

(a) C, against a.

Figure 13.- The lor@tu&iml charecteristics of the model equippea with various inlets.

(b) C, against CL.

Figure 13. - Continued.

/.2

l.0

.8

.G

.4

.2

0

-. 2

"4

CL

"

-4 0 4 8 /2 I6 20 24 28 32 0 0 0 0 0 0 0 4 8 I 2 1 6 2 0 2 4 0 ._ A V 1,. d a

G deg

( c) c T ~ against a.

28 32

Figure 13 .- Continued.

12

LO

.8

.6

.4

.2

0

9

-.4

I CL

o .I .2 .3 F? .5 .6 P

(a) cI; against cD.

Figure 1-3 .- Concluded.

,

(a) C, zgailzst a.

Figsre 14.- The longitudinal characteristics of the nodel equipyed with various inlets and ving fences.

-.04

-.08

-. 12

-. 16

(b) CE against CL.

Figure 14.- Contbued. .-

I I

A2

.8

.6

.4

.2

0

-. 2

CL

-4 .. -4 0 4 8 I2 I6 20 24 28 32

(c) cL against a.

Figure 14. - ConLinued.

.8

.6

1 .2

0

”2

74

(d) CL against CD.

Figure 14. - Concluded.

./2

.08

.04

0

-.04

Acm

./6

-. /2

.08

-.04

""_

0 4 8 /2 /6 20 24 28 32 e, del7

Figure 15.- The deviation vith angle of zttack of the pitching-momect coe3ficient from (dCddu),=-, for the model equipped x5th the pro- ductLon t a i l and various i n l e t s .

"04

. 0 8

.cw

Acm o

-.04

-08

.04

0

-.04 Acm

-.m - t4 0 4 8 1.2 16 20 24 28 32 e, &?

Figure 16.- The deviation with angle of attack of the pitching-moment coefficient fYom (dCdcia)a=o fo r the model equipped with the Fro-. duction tail and with various inlets and wing fences.

./6

./2

.08

.04

I-! 0

-.04

-.OB

-.I2

-.I6

-A0

-24

-28 -0

-.m -.08

-.I2

-.I6

-20

- 24

(a) b2ets off, C, against a.

1

79

Figure 17.- Effect of horizontal-tail col?figuration on the longitudinal characteristics of the model equipped w i t h various inlets.

./6

. /2

.04

00

-.08

-./2 n 0

-.16

-5 0 -.20 cm

-. 24

-.08

-. 20

(a) Continued. Inlets off, Cm against CL.

Figure 17.- Continued.

.

LO

.8

.6

.4

.2

0

-2

- .4

CL

-4 0 4 8 /2 /6 20 24 28 32 " 0 0 0 4 8 /2 /6 20 24 28 32

0 '. e, deg

(a) Continued. 1nle.l;~ orf, cL against a.


LO

.8

.6

GL 2

- 1 -2 0 -.4

.3 .4 -5 .6

(EL) Concluded. Inlets off, CL against; CD.

F i v e 1.7. - Continued.

.

.08

.04

”0

-.04

“08

-. 12 LO

-. /2

-24 -

(b) In l e t h,, C, against a.

Figure 1-7.- Continued.

LO

.8

.6

I Ci -2

0

-.2

-4 -4 0 - 4 8 1.2 16 20 24 28 32

V 0 0 4 8 12 16 20 24 28 32 L ._

deq

(b) Continued. Irilet 113, CL against a.

Figure 17.- Cmtinued.

I .

.08

-04

*' 0

-.04

-.08

-. 12

' m -./6

.-i 0 -po

"04 724

"08 72'8

-./2

-. 16

720

-. 24

00

-.4 ,2 0 .2 i4 . 6 .8 LO CL

(b) Contimed. W e t "3, C, agzinst CL.


.2 -3 .5 .6

(b) Concluded. Inl.et 03, CI, against CD.

Figure 17 .- Continued. L l

. ( c) Inlet D2, Cm against a.

Figure 1-7.- Coctinued.

NACA RM LfihS17

-.4 7 2 0 .2 .4 . 6 .8 LO

CL

(c) Continued. Inlet D2, C, agzinst CL.

F i w e i7 .- Continued.

.8

.6

0

-. 2

"4

_.

I

L

h3 I N

. _ -4 0 4 8 12 16 20 24 28 32

a 0 0 4 8 12 16 20 24 28 32 Kl a

CG deg

(e) Continued. 1nl.e.t D2, CI; against a.


w 0 IIQ

.8

.6

I .2

0

0 .I .2 .3 4 .5 .6 9

(c) Concluded. Inlet D2, CL against CD.

Figure 17 .- Continued .

H

I .

./2

.04

-. 04

-.08

-. 12 Q O

-.04 724

-. 08 -. 28

4 0 4 8 12 16 20 24 28 32 G deg

(a) Inlet Dl, C, agzinst a.

F i w e 17 .- Contimed .

-

I I

-4 0 4 8 I2 I6 20 24 28 32 a 0 0 4 8 12 I6 20 24 28 32

0 0 G des

(a) Continued. Inlet Dl, CL against a.


LO

.8

.6

.4

.2

0

9

-.4

cL

0 .I 2 .3 4 .5 .6 J 0 0 ./ .2 .3 .4 .5 6

0 c cD

(d) Concluded. Inlet Dl , CL against CD . Figure 1'7 .- Continue3.

-08

.04

- .04

-. 16

-.OB


.I2

.08

.04

G 0

0.04

-.OB

Gm -42

-.It$

1': 0 -20

~04

4 2

0.08

=. 12

(e) Continued. I n l e t Do, C, against C-, . Figme 17.- Continueci.

- + w

I I I N

LO

.8

.6

I .2 0 i I ( e ) Continued. Inlet Do., CL agsinst a.

1.0

.8

.2

0

(e) Concluded. I n l e t Do, CL against C,.

Figure 1‘7. - Concluded.

I

I

.2

0

-2

-.4

,4

.P

0

-.. -.4

.8

-6

4

.2

0

-2

- .4

-. - rlgure 18.- The vzriation of dw&& with lift coefficient for %?e model equipped with vmious horizoztal-tail arrangenents and uith and without various L n l e t s .

.8

.6

.4

.2

0

-2

-4

NACA RM L54517

COPT-garation hie:

.6

.2

0


102

08

.04

0

-.04

-.08

- ./2

-./6 0

G 7 -20

-. 2#

0 -28

-04

- 08

- ./2

- ./6

-20

(a) Continued. Fla3s neutral, Cm against CL.

Figure 19. - Continsled.

- "

I . 1

L 2

LO

.8

.6

0

-.2

-.4 -4 0 4 8 12 /6 20 24 28 32

0 0 4 8 /2 16 20 24 28 32 a=, deg

(a) Continued. Flaps neutral, CL against a,. Figure 19. - Continued.

w

12

LO

.8

.6

4

.2

0

-. 2

- .4

CJ L

o ./ .Z .3 4 .5 .6 0 1 0 ./ .2 .3 .4 .5 .6 .7

CO

(a) Concluded. Flaps neutral, % against CD.


, .

.04

0

-.OB

- ./2

0 -.I6

0

-.O#

-.08

-. /2

-./6 -4 0 4 8 12 16 20 24 28 32

a=, deg

(b) Flaps deflected, C, agahst a.

Figme 19 .- Contimed.

I" 0 cn

112

LO

8

.6

.4

.z

CL

n -4 0 4 8 I2 /6 20 24 28 32

LI

(b) Continued. Flaps deflected., CL against; a.

Figure 19.- Continued. FJ a

. /6

./2

.08

04

0

-.04

-.08

- ./2

0 -.I6

-20

cm -24

0

- . 0 4

-. 08

-./z -. /6

(b) Continued. F l q s deflectec, C, agzinst CL.

Figce 19. - Continued.

li2

LO

.8

.6

.4

.2

n

CL

'D 0 ./ .2 .3 4 .5 .6 .7

(b) Concluded. Flaps deflected, CL against CD . Figure 19. - Concl-uded.

F c1 *

. .

.2

0

-2

-.4

-.6

-.8

. 7 4 -2 0 .2 .4 .6 .8 LO L 2 CL

F i v e 20.- Variation of dCddQ, w i t h l i f t coefficient f o r the model equi-oped with inlet DL o r Do, horizontal t a i l T . z , and various

favorzble -wing coEf igurations .

! 10

(a) CL and C, against a.

Figure 21.- The longitudLr-al charecteristics of the model eguipped xitin inlet Dg, flight fences, leading-edge modificatim, and production or drooped tail. Trailing-edge flzFs zeutral.

.6

.2

./

0

(b) CD an& C, against C-L. Figure 21 .- Concluded.

6 2

.2

0

NACA X! L5kBi7

(a) CL and C, against a.

Figure 22.- The longitudinal characceristics of the model equipped with inlet Dg, flight fer-ces, leaaing-edge mdification, an5 production or &coped tail. Trailing-edge flaps deflected 403.

.7

.6

.5

.4

.3

.2

./

0

c,

.08

. 0 4

0

-. 04

-.08

- ./2

-.I6

"20

cm

(b) CD an6 C, against CL.


1 14

.2

-. 2

- .6

.4

2

0 G?? q "2

- .4

- .6 -.4 ,3 0 .2 .4 . 6 . 8 LO LZ

CL

Figwe 23.- The variatioll of &C,/dCl with l i f t coeff ic ient for the riodel equisped with i n l e t Do, flight fences, leading-edge modifi- ca-kior,, and the production or drooped t a i l . -

W C A R3l L54B1-7 11 j

-4 0 4 8 /2 /6 20 24 28 32 e, deg

(EL) C, against a.

.- ETfect cf external stores on the longitudinal c h ? e c t e r model equipsed v i th vwiaus tails in inlet conTigwations

istics .

l l b

./2

.08

.04

uo 0

-.04

-08

-. /2

-. 20

cm -.24

-. 28

L V 0 w.32

-04

-. /6

(b) C, against CL.


.

12

.8

.6

4

.2

0

-. 2

- -4

CL *

. . -4 0 4 8 /2 /6 20 24 28 32

(c) CI; against a,. Figure 24.. - ConLinued .

12

LO

.8

.6

.4

.2

0

-.2

CL

.I .2 .3 .4 .5 .6 .7

.8

.6

0

-. 2

-.4 .08

.04

0

-.04

-.08 cm

-. /6

-. 24

(a) CL a16 cn? egainst a.

Figure 25.- EZfect or" inlet mass-flov ratio on tbe longitudinal chmac- ter is t ics of the model equipped with i n l e t DL and horizontal teil T . 2 .

120

.6

.5

0 .08

.04

0

-.OB

-. /2 Gm

-. /6

-. 20

-28 -

NACA

.4 -.2 0 .2 .4 .6 .B LO 12 CL

(b) CD and C, agafns-l CL..

c

6Z NACA EIM L54B17

/. 0

.8

.6

.4

8 -4 0 4 8 /2 /6 20 24 28 32 dep

(a) CL end C, against a.

Figure 26.- Longst-dim1 znd lateral-control characteristics of the model equisped wFth an outboard aileron. Configuration A + Dl f T.Z.

.6

.5

.4

. /

0

(b) CD and C, zgainst CL.


. -02

. o/

-02 >cf+ction, deg

(c) C, m d C2 against CL. Figure 26.- Concluded.

124

12

1.0

.8

.6

.4

.2

0

-.2

". 4

7 6

cL

./6

./2

.08

.04

0

7 04

7 08

T / 2

-:/ 6

T.20

cm

-

(EL) CL and C, against a.

Figure 27.- Longitudinal s?nd lateral-control chzracteristics of the model equipyed with an outboard aileron. Configwation A + Do + T . 3 + '0.306 (0.652 - 0.958) + flight fences. -

.7

.6

.5

.4

.2

0

(b) % and C, against CL. Figure 27.- Continued.

126

-. 02

.03

.02

f 02

703 -

34ect ioz, 9sg (Le& aileron;

8 -4 0 4 8 12 16 20 24 28 32 deg

Figure 27 .- Concluded.

.

(a> CL and Cm against a.

Figwe 28.- Longitudinal znd lateral-control chracteristics of the mdel equLpped xith a full-span zileron. Configuratior- A + Do + Tam + I,.30~(0.652 - 0.958) + P l i g h t fences.

128

.7

.6

.5

.4

.3

(b) % and C, against CL.

Figure 28. - Continued. -

.

.

(c) Cn and C2 agzinst a.


/. 2

/. 0

.8

0

.

(a) CL and C;, against a.

F i v e 29.- Longitudinal and latersl-control characteristics of the ntodel equisped with differentially deflected flaps and outboard ailerons. Configuration A + Do + T . 3 + 10.306(0.652 - 0.958) + flight fences.

.7

.6

.2

0

. /2

-04

-./ 6

(b) CD and Cm against CL.

Figsre 29 .- Co-n-tinued.

->

( c ) C, and c2 =gainst a.

.

Figme 29. - ConcIuCed .

/.2

LO

.8

.6

.4

-2

0

CL

Figure 30.- Longitudinal and lateral-control characteristics of the model equipped with an outboard aileron. Configuration A + Do + T , X + X,.,5(0.708 - 0.958).

.7

.6

.2

. /

-/6

. /2

.08

-/6

NACA RM L54317

0 .2 .4 .6 .8 LO 12 CI 1

i

(b) cD ana C, against CL.

FLgure 33.- Contizued.

.

Deflection, aeg (Left dleron) (Raps)

a 18 a 9

40 40

D 3 40

P -9 40 A -18 40

a -3 40

(c) C, and CZ against a.

Figure 30.- Concluiied.

"

1-36

/. 2

.8

.4

.P

0

(a) ana C, against a.

Figare 31.- LongitucTinal m d lateral-control chaxacteristics of the model equipFed wCth an outboard aileron. Cofifiguration A f Do + E0.15(0.70&/2 to 0.85&/2) -I- Sf = &03. -

.7

.6

.5

.4

.3

.2 co

./

. /2

.08

5 /2

0 .2 .4 .6 .8 /.U 1.2 ci 1

(b) CD end Cm zgainst CL.

Figwe 31. - Continued.

(Len dIercr) (rape) &.flect!or., deg

0 0 40 18

A -1s 40 40

02

.o/

.02

.o/

-.03

NACA RM L54Bl.7


c

.6

.4

CL .2

0

-. 2

-.4

. /2

.08

.04

,20

(a) cL ana cn against a,.

Figure 32.- Lorrgitudinal acd la<eral-control character is t ics or’ the eqaipped u i th sol id Tlap-type spoilers. Configuration A f Do c io,30~(0.652 - 0.958) i Tlight fences.

.6

.4

. /

. /2

.04

-. /2

-. /6

CL

(b) Cy) and Cm zgainst €$,. Figure 32.- Continued. -

. 01

-. 02

. o/

0

-. 02 -. 03

Spoiler deflection, deg

0 4.7 0 9.35 A 13 v 45 D 55 a m 1

i

( C) Cn m-d C 2 zgainst a.


14 2

/.2

/. 0

.8

.6

.4

.2

0

-. 2

-:4

-: 6 . /2

.08

.04

0

: 04

5 0 8

:/2

-. /6

CL

cm

.

(a) CL an2 C~ against a.

Figure 33.- Longitu&inal an2 lateral-control characteristic& or' the model equiqed vi th perforated flap-type spotlers. Configuration A + ilo + T.Z + rC . 3 0 6 ( ~ A52 - 0.958) + nigqt fences.

.

.6

.4

.2

. I

0

-.6 -.4 -.2 0 .ZC, .4 .6 .8 1.0 1.2 L

(b) CD and Cm against CL . Figure 33 .- Continued.

143

.o/

"

8 4 0 4 8 12 /6 20 24 28 32 E* deg

( c ) cn an& C2 agairst u . Figure 33 .- Concluded.

.o/

-.o /

. o/

-03

Figure 3b.- Variations of the yew and roll characteristics of the model hith spoiler defLection. Configmation A +- Do + T . 3 f

Io.g06(0.652 - 0.958) + flight fences.

A

.8

.P .6

.4

Hi? - Po q0

.2

I-

o\ .!=

0 - % - 4 4 8 I 2 16 20 24 28 32

a, deg

(a) Exit, NU open.

v; VO q 0

Figure 35.- Variations of - ana IEe - with mg1.e 03 attack fo r

the model equipped with varims inlets. R = 9 x lo6.

. 1

- I a

- --

.6

v0 .2

0

l.0

8

,6

A

.2

0

-8 - 4

1 4

(b) Exit, half open.


0 4 8 f2 /6 20 24 28 32

E, de9

.8

.6

.8

.2

.6

0

0

(a) kit, IXU open.

Figure 36 .- Variation of - arid Vi 11, - Po vo s,

wlch angle or' attack for the

model equipped wiLh inlet Dl and horizontal kil T -. R = 9 x lo5. .2a

z * 0 >

F 9 Y

*r D n

4/i "0

.4

-2

0

1.0

.8

. .6

He - Po (70

.4

.2

0 -8 -4 0 4 8 /2 16 20 24 28 32

a, deg

(b) Exit, 39 percent open.

Figure 36 .- Concluded,

I

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