RESEARCH MEMORANDUM - NASA · RESEARCH MEMORANDUM INVESTIGATION AT HIGH SUBSONIC SPEEDS OF THE...

RESEARCH MEMORANDUM

INVESTIGATION AT HIGH SUBSONIC SPEEDS OF THE STATIC

LONGITUDINAL AND LATERAL STABIIJTY CHARACTERISTICS

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OF TWO CANARD AIRPLANE CONFIGURATIONS t

By William C. Sleenian, Jr.

Langley Aeronautical Laboratory 2.' Langley Field, Va.

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is 8"

= . :q NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

WASHINGTON December 3, 1957

Restriction/Classification Cancelled

https://ntrs.nasa.gov/search.jsp?R=20050019253 2020-03-03T00:08:05+00:00Z

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NACA RFI ~ 5 7 ~ 0 8

I?!~TVESTIC-P?TION AT HIGH SUBSONIC SPEZDS OF THE STATIC

O F TWO CANARD ARPI-JLNE COIWIGiiEwTIONS

By William C. Sleeman, Jr.

The present irwestigation w a s conducted i n the Lmgley high-speed 7- by 10-foot t m e l t o d e t e r d n e the s ta t ic longi tudina l and lateral stebil- i t y cha rac t e r i s t i c s et high subsonhc speeds of two canard airplane con- figurations previously tested a t supersonic speeds. The %ch nmber range of this icvestigation extended from 0.60 t o 0.94 ea& a I ; l u r i m m angle-of- attack range of -2O t o 2k0 was obtained at the lmest test Mach number. Two wTng plan forms of e q w l area were studied i n the present tes ts ; oce w a s a 60° de l t a k-ing and the other was a trapezoid wing hzving en espect r a t i o of 3, t ape r r a t io of 0.143, and EI-I unswept SO-percect-chord =ne. The cmard control had e trapezoicial plan form and i ts area was approxi- mtely 11.5 percent of the vir?? area. The m o d e l also hed a low-aspect- ra t io highly swept ve r t i ca l t a i l azla twin ventral f ins.

The longitudinal control chmacterist ics of the mdels were consist- ent w i t h past experience a t l o w speed on cmard configurat ions in %hat stallilzg 03 the caznsd surfece occurred at moderate and high control Oeflections for notierate values of angle of attack. This s t a l l i n g could impose appreciable limitations on the nzxirnun trim-lir ' t coefficient a t ta inable . The control effectiveness and maxinun value of t r i m l i f t was significactly increased by zddition of a body f h p having a conical shape and located s l ignt ly behip-d the carlard surface on the bottom of the body.

Addition 03 t h e canard surfece a t 0' deflection had r e l a t ive ly l i t t l e e f fec t on overa l l d i rec t iona l s tab i l i ty of the delta-wing con-pigur&tion; however, deflection of the canard surface from Oo t o loo hed E. lmge favor- able effect 02 di rec t ione l s tab i l i ty a t high zngles of a t teck f o r both the trapezoid- m d iielta-wing configurations.

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?resent in te res t in canard airplane configurations arises frm- ;potential perfomance benefits possible a t sGgersonic speeds i n cozrspari- son t o conver;tional tail-reamard an6 tailless airplane designs. The canard arrangeaer,t also offers a solu-kion to p rob lem of balance which are becodng nore aczte wit'n the design trend towzrti =are rearwzrd engine p1acercer.t End accompqying rezrward center-of-gravity Position, partic- ularly for mltiengine arrangerients. Tne nain problez encomtered in the pas t for cenzrl! configzrations was associatee with s t a l l i ng of the canard s.mface vhich severely l i d t e d the alloxable center-of-gravity t r ave l an5 -cke m m i m u n trix l i f t of K?e airplzne . (See ref. 1. ) This problem as well as tne 5ireetional stability d i f f i cd l t i e s fotrnO for some canard configurations icdicated the existence of sorce formidable sub- sonic problem, and the potent ia l rewards t o be gdned a t subsor;lc speeds by Lsing canard controls did n o t n e r i t s o h t i o n of these problem. The aforementioF-ed supersonic performace ber-efits offer a? effective stimulus to reneved e f fo r t an I"inbi??-g solutions for t i e known lov-speed groblem agd for emloring the general aeroQnmic characteristics of cenzrd air- plane configurations a t susersonic speeds.

Arr e-erirr-erital study has been con&xted a t su2ersonic speeds of sone generalized airplane cor-figurations which use canard surfaces for longitudinal control. Soxe of the resu l t s of khis strzdy are presented i n reference 2 and include longitudiml and lateral stabil i ty character- i s t i c s obtained a t Mach numbers of 1.41 and 2.01 for two canard airplane configzat ions &vir-g a 60° delta wing a-6 an aspect-ratio-3 trapezoid wing. Tce coafigurations of reference 2 were selected on the basis of available Fr-formtion as a concept of a gooa supersonic airplane configu- ra t ion heving mhinux changes i n aerodynazdc chwacter i s t ics ikon sub- sonic to susersonic speeds. It woLzid therefore aypem desirable-in obtaining subsonic infoxation on &n advanced generalized cacard xodel t o use the seme conf igwat ims which were tested a t supersonic sseecs.

We present investigatim was conEucted i n the Langley high-speed 7- by 10-foot tunnel w i t h the mdels of reference 2. Longitudinal and l a t e ra l s t ab i l i t y cha rac t e r i s t i c s vith and without the canard surr'ace aeflecteci were abtained over e Mach number range fron 0.60 t o 0.94 en6 a aax iu -mgle -o f -a t t ack range fron approximtely -2O t o 2k.O a t t h e lowest t e s t Xkci number. In aCdition t o tests of the complete model configurations, assorted 'DreaMown tests were Z S e t o d e t e d n e e f f e c t s of addition of canard surfaces, the vertical tail, ma the ventral f ins . Sone br ie f tests were =de w i t h a bow f l ap &n12 a canard flap i n a t t e q t s to increase the maximam angle of a t tack for which the model could be trimd in p i tch .

NACA RII ~ 5 7 ~ 0 8 3

Lzter.11 s t a b i l i t y r e s u l t s of this imrestigation ere referred t o the bcdy axis system whick i s shown i n figure 1 together w i t h an indic&ion of positive Girections of forces, noxents, and =gular deflections of the Eodel. The l i f t and &rag character is t ics :resented a t zero sideslip are, respectively, norm1 zr?d p a r a l l e l t o the re la t ive wind as shown i n the side viev or" t l e model in f i g m e 1. Moxect coeff ic iects are given about the reference ceDter shown i n figure 2 (located on the body center lice a t body s ta t ion 25). This position corresponds to a location 17.8 per- cent nem a e r o d y n d c chard ahead E d 7.75 gercent m e a aeroaynunic chord behind the leading edge of the nean z e r o d y n d c chord f o r the trapezoid wing and de l ta wing, respectively.

pitching-moment coefficiect , Pitching moment qSE

Cl rolling-xomext coezficient, Rolling moment

G b

cn yawing-moment coefficient, Y&.-wing m m n t

qSb

cy la teral-force coeff ic ient , Leteral force ss

? air density, slugs/cu f t

S wing area (including aree inside body), sa_ f t

b wing span, ft

4 NACA RI~I ~ 5 7 ~ 0 8

a argle of a t teck of fuselage center line, deg a

F angle cl' sideslip, deg

SC de f l ec t io l mg le of cmarc surface (poEitive wi th t r a i l i n g edge do-m), deg

Subscripts :

9 denotes gartial derivative of a coefficient w i t h respect t o s i d e s l b ; f o r examgle, Cnp = acn

t denotes effect due to addi t ion or' ver t ica l t a i l

M O D 3 D E S C R E T I O N

The basic model configurations tested ere shown in f igure 2 with SOTE of the perticent nodel dimensions. Details of the model geometry are given i n table I and coordinates or' the body are given as table 11.

Two low-asgect-retio-Xing glm f o r m of current interest were tested Y

OE t he s a e boEy t o study t i e character is t ics of two cmard airplane con- figurations having a highly swept de l ta wing and a trapezoid wing of l o w s-xeep. The 60' Celt& wing haE an Etsgect ratio of 2.31 and g. raxCnun thick- cess of 4 Fercent of the wing chord. The trapezoid wing hac5 an aspect r a t i o of 3, t ape r r s t i o of O.lb.3, and an mswept 80-percent-chord l i ne (28.82O sveepback af .Yge quarter-chord line). The maxirmm thickness of t?"e trapezoid wing was 4 percent of the wing chord also. Both of t i e wings tested were mcie or' s t e e l and had hexagonal a i r fo i l sec t ions with

edge. seniasex argles n o r d t o the leading edge act2 n o m 1 t o t h e trailing

The canara control surface which hzd hexagonal a i r fo i l sec t ions (see f ig . 3) of the same descrl2Cion as for tine wings w a s remotely controlled over a def lect ion rmge frolr Oo t o 15".

TITO a1iri l iary devices (see f ig. 3) were studied as possible meam for obtaining positive pitching-mment irrcrexents t o aid the canard sur- face is. providing t r i d 9 - g moxerks a t kigh angles or" attack. The body nose f l ap -was a conical segnent 'having a radiirs equal t o the body radius

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a t the f lap leading edge. ?&e canard f l a p was a weOge which sinulated a s p l i t Flag deflected TO0.

The node1 w i t h the trapezoid w i n g w s s modified for part of t'ce investigztion by adding a 5-inch-long cylir_drical extension to the bzse of t h e o r i g i n a l b o u . (See f i g . 2.)

Tests

The present investigation was conducted In the Lmgley high-speed 7- by 10-foot tunnel over e. ,Mach number reage from 0.60 t o 0.94. The everage t e s t Reynolas n-mher based 0x1 the mean zerodyn&c chord w a s agprouimtely k . 5 x 10 f o r t h e delta 'ding and 3 .3 x 10 for the trap- ezoid xing a t 31 = 0 .go.

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The nodelwcs mounted on a six-cmgonent internal strain-gage balance of the sa.^ descriptioc as t h a t used i n t h e tests of reference 2 end the belence was supported by a variable-angle sting. Longitudinal stability c k r a c t e r i s t i c s were investigated a t zero sideslip thoughost the angle- of-attack range with the Mech number and canard deflection held constant. La tera l s tzb l l i ty der iva t ives of th i s inves t iga t ion were obtained from tests conducted tkrough the angle-of-attack raage with the mo6el a t fixed sideslip angles of *bo. The naxirnm range of anzle of a t tack extended from approximtely -2O t o 2 4 O a t the lovest test Mach number. Some Uni t ed tests were cor-ducted t'nxough a rmge of sideslip aagles from -bo t o l 2 O w i t h the mdel angle of' at-lecb held constant.

Corrections

Jet-bour-dary correc-lions to t he ang le s of a t tack 2nd drag coefzici- eDts de t edced f rou r e fe rence 3 were added t o t'ne data. Blockege cor- rections zpplied to t'ne Mach r"Ders were Oeternined from reTerence 4. Drag coefficients have been corre.cted f o r a srnall tiznnel-buoyancy effect acii corrections have also been applied t o the drag coefficients suc3 that the base-pressure cocdltions correspond t o free-stream static pressure.

The ar-gles of a t tzck an& s ides l ip of the model have been corrected for defiectiorz OF the balance acd stlng izzder load.

3ACA RM 1.57~08

The effect of the canard on the aerodynaxic charac te r i s t ics in p i tch 5

oI' the delta-xing mdel is presented in f i g a e 4. Effects of cmard deflection on the model wi th the delta a d trzpezoid wings are giver- i n figures 5 azd 6, respectively, and ef fec ts of a b d y nose f lap md canard f l a p a r e s'nom in f i gu re f o r the rrodel w i t h the iielta wing. The e f fec ts of the vent ra l f ins 03 l ong i tud ina l cbac te r i s t i c s fo r t he iielta-wing con2igur&,tion are shorn i n figure 8, an2 ef fec ts of tlrle 5-ir,ch-&fter%ody extension for tbe node1 wit'n the trapezoid wing are show- in f i gu re 9.

Lateral s tabi l i ty der ivat ives of tke model witn the deltz wing which show ef fec ts of the cmard swface, canard deflection, and ve r t i ca l t a i l &re presented. in figures 10 aii 11. Figure 12 ?resents the effects of both the v e r t i c a l t a i l enE the ventral f ins on the l a t e r a l s t e b i l i t y derivazives or' the trapezoi6-uing configpration. AerodynarnLc cimracter- i s t i c s i n s i d e s l i ? showing ef fec ts of the canard sLLrI?ace and canard deflection are presenteC in figsre 13 fo r t'ce &elta-wing cocfiguration.

The v a r i a t i m w i t h Mach nuzber of the longitudinal and direct ional stakili-by a t zero l i f t , of dninun drag coefficient, ar-d of maximan lift- i k a g rzt.io i s gresented in l i gxe 14. LongituZinal chracter is t ics for trin conditions tzroughaut the lift range are given i n figuse 15. Pitcking-monent results are presented in f igure 16 an6 these r e su l t s show ef fec ts of s t a t i c mg2 .3 on t he con t ro l ch rac t e r i s t i c s of the Eodels. Increnental effects or" the canard surface Etnd canard deflection on l a t e ra l s t aa i i i t y p snmete r s fo r t he delta-wing model are presented in f igure 17. Sffects of t'ce verrtral fins on overal l d i rect ional stzbil- i t y and ef fec ts of canard dePlection on the t a i l contribution t o direc- t i o z a l s t a b i l i t y a r e s:?own i n f i g w e 18.

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DISCLESION

I? Cesirable airplzr-e arrengexent would heve the center of gravity located such that tlle fu7-l l i f t i ng cepab i l i t i e s of the wing can be rea l ized in f l igh t with the a i rp lane t r imed et the most forward posit ion of the center of gravity end be no less than neutrally stable with the center of" gravity i n the r e m o s t p o s i t i o n . The cmarci airplane zrrange- nent therefore should typically have a nore forward center-of-gravity position thzn tailless or tail-rearward eirglanes BS r e su l t of the 6estabilizilzg contribution of the canard surface. For exaqle , addi t ion of %he camrd surface to the present delta-wing configuratfon caused a destabi l iz ing s ta t ic nzrgin shift of ap2roximxxtel.y 10 Dercer-t E ( f ig . 4) v through the test Y!ch nmber range. The rather foriard location 03 the reference cezter of gravity for Yle present mdels colnciCed w i t h the

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Eoment center used in the results Tor refereme 2 and does r,ot neces- sarily indicate an optimum center-of-gravity location which would depelld. on both longitudinzl and d i r e c t i o m l stebility and control. T'ne r e su l t s smmarizeb in figure lk show that t'ce low-l i f t s ta t ic =gin was quite large for both wi,ng slan f o r m at a Mach nurdber of 0.60 md increased =bout 3 percent 'c es the bbch number increzsed t o 0.94. The high level of lozgi tudina l s teb i l i ty sham in figure 14 would be excessive for an aircrar't designed t o f l y a t supersonic as well as subsonic speeds and the rroment reference would heve t o be sh i f ted rearw&,rd for appl ics t ion of the present data t o such an arrmgement.

Loagitudinal Stability C"-Ccteristics

Conkrol effectiveness of the basic models.- Effects of canard deflec- t i o n on the longitudinal stability c h r a c t e r i s t i c s of the nodel with the del-La and trapezoid wings are p resen te s i n figures 5 m d 6, respectively, and tria longitudinal chexecteristics obtained from figures 5 and 6 are s w m r i z e d i n figme 13 f o r the lowest and highest test Mach numbers. The r e su l t s of figure 15 show tha t the n a x i m u n lift coefficient for which the n d e l could be t r i m e d with Tie cenzrd surface deflected l5O w z s a l i t t l e greater than 0.30 a% M = 0.60 for both wing p l m f o r u s and decreased to approximte ly 0.20 a t M = 0.94.

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P Reasons f o r t he low values of naximun t r i m lirt coeEicients obtained a re qpa ren t f ron the pitching-moment curves of figures 5 and 6 which show a re la t ive ly high static m g i n and the occurrence of s t a l l i n g of the cmard surfece at the highest deflection angles. In the ebsence of bow and wing induced upwash, the cward angle of a t teck vould be ecykL t o the sirplzne angle of a t tack plus the deflection angle, aGd, therefore, a t high initiel defleckions, stalling of the canard surface would be q e c t e d t o occur at r e l a t ive ly low e i rp lme angles of attack.

In o rde r t o assess the canmd control character is t ics for mre rea- soneble values os" s t a t i c -gin, the h t a of figures 5 and 6 have been referred t o different moment-center locations an6 those results are given i n f i g u r e 16. T??e pitching moments presented i n figare 16 are f o r low- lift s t c t i c Illargios of 5 and 10 percent or' the nean aerodymm5.c chord fo r 6, = Oo. These r e su l t s shov that the r r a x i m t r i m lift vas exbended by reciucing the s t ab i l i t y ; however, there m s very l i t t l e control effective- ness cbove a l i f t coefficient of approximately 0.7 fo r either the delta- wing or the trapezoid-wing Eodel. The problem of the minun airplm-e t r i m lift be iw l imi t ed by s t a l l i n g of the canard was not encountered i n the tests at supe66onic sceeds" ( r e f . 2) because of ' the ber-eficial effects of susersonic Kich number on m?irmm lift, c k z a c t e r i s t i c s . It appears, thererore, that f o r e configuretion having a smll t r a s o n i c aer0Q-c- center shift, the problem of canard c o ~ t r o l e r f e c t i v e n e s s will l i e sri- w i l y i n the subsonic region rather than supersonic.

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A comparison of pitching-moneni; resuLts far t:ne delta-wir!! and trapezoid-whg models >resented in f igures 5 and 6 S~OTV-S some differences which skc'clld be explzined. The values of pitching noxent a t zero l i f t ( f ig s . 5 EEL 6) shm t h a t a larger pitchiag-xornent coefficient ves pro- duced by a given cmtrol i ief lect ion on %he trapezo2.d-ving rcodel tha~~ for the delta-wing model. $,%en interference i s neglecteL, the camrd sur- face s'r_oul& be expected t o provide the sarce moment regardless of which wing i s 3ehin2 it, and t h i s mment shodd trir- the cozfigJre.tion having t i e lowest sta-Cic =gin t o <he k-igher t r i m - l i e mlue. Tne two wing plan Zorms had the SECT^ area; kowever, the man aerodymdc chcrd for the C e l t z . wing vas Epprosimtely 26 percent greater t:mn tha t for the trap- ezoiC wing, and C3is differezce i s ref lected in the differences ir. canard effectiveness at zero l i f t &n-d in the camrd cor-tribution t o longitudinel s+,abiiity. No% all the differences iz control effectiveness md rraxinun trix lift can be at t r ibuted t o differences in the reference len&n used ir_ the coefficlents; hawever, mxt of the differences shown for the two wing plan foms can be at t r ibuted to t3is source.

Auxiliwy control devices.- Several xeens h v e been s tudied in the gast for increasing the l i f t ing capabi l i t ies of canzrd scrfaces a t low speeEs such as acdition of leadill@;-edge slats acd various ty-pes of trailing-edge flzps. (See r e f . 1.) These devices EZLC recent; clevelop- ments +n E-ie use of bolmd&ry-lz.yer cantrol offer promising means for rzaSeris l ly iacresskg the l i f t ing casa 'oi l i t ies of a canard surface at low speeds. Another a-psroach to the problen of at ta ining t r i m a t nigh L i f t coe4ficier-ts is that of re l ieving the canard surface of the bulk of the xonents t o be trim& by use of' auxiliary trim??zg devices. Some l i r l i t e3 t e s t r e su l t s were o-atained w k t h E body f l a p (shown i n f i g . 3) used ES a device t o provide a posit ive pitchFng-monent increment. Test resl;l-Ls obtained x i t h this f lap for t he delta-wing nodel ( f ig . 7) at e. hhch m&er of 0.60 show khat an appreciable increlzent in pitching- rcornent coefzicient resulted from the addition of the f lap . Eear zero l i f t , -LIIFs increxellt we8 apsroximtely equal to the increznent which would 5 e obtaineC wi th the basic canerd surfsce Oeflected 7". The mi- 111m l i f t coe2ficient st which the model coi;l& be trimmed w i t h a canard slxface deflected 10' w&s increased frm 0.30 t o 0.55 by addLtion of the body flap. A benefit of the bo- f lap, in edc i t ion t o the besic incre- Tent a+, zero lift, wzs the increase in control efzectiveness at high l i f t coefficients. The drag increTer:t of the body f l a p shun i n flgure 7 szggests that it would be suFted only 2s an zis! i n trimming fo r Low- speed f l i gh t .

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I n addikion t o the body fla?, sore r e su l t s were obtained u i t h a YO0 wdge attached t o the czmrd surface as shm- i n figure 3 which simu- la ted a spl i t f1.q derlected 300. Addition of this sinLLa.ted cmard f lap provided mprecia'cle gains i n t r i m l i f t coeff ic ient ( f ig . 7) over the I

besic arrangement; however, t h i s f lzg w a s nuch less e f fec t ive than the . bciy f lap. Of course Ir-uch more effective carard f lap 3sranger:ents a re

d

NACA RM L57J08 9

SossibLe, such as e single or double s lo t ted flap; however, a l l these devices loed tfie canard more heavily rather than provide z t r i m i n g noment tha t would be independent of the canard.

Effect of vez t r a l f i n s and extended afterbody.- Effects of the ven- t r a l f i n s on the longitudillel characterist ics et the lowest and higbest t e s t Yach nurzber are given in f igure 8 w-d show essent ia l ly no e l f e c t of tb-e vent ra l f ins on the longi tudina l s tab i l i ty of the delta-wing model. The lack of asprecisble effects of the vent ra l f ins on either longitudi- na l or d i rec t iona l s t ab i l i t y ( f i g . 18) probably results 2rom the f a c t thzt the center of area of the vent ra l f ins i s loczted. close t o the mment reference.

Tests were -de w i t h an extended afterbody on the trapezoid-wing configdrstion i n order t o es tabl ish the subsonic s tabi l i ty level for st?l evaluation of e f f ec t s of afterbody extensioz on t'ne aerodynamic-center s h i f t i n going from subsonic t o supersonic speeds. The ?resent tests showed essent ia l ly no e f fec t of the afterbody extension 02 longitudinal sta'oility a t low and moderste l i f t coefficients. (See f i g . 9. )

Lzteral Stabi l i ty Character is t ics

Effects of canerd swface snd cznard deflectioc on laterel s t a b i l i t y derivatives.- ETfects of addition of the canard surface and canard deflec- t i o n on t h e s t a t i c h t e ra l stability derivatives of the delta-wing nodel are presented i n figure 10 and ir-crenents obtained from 2i-e l0-sre s-amzrized i n f i g u r e 17 f o r a Mach nm-ber of 0.60. These r e s u l t s show very l i t t l e e f fec t on 6 i r ec t iona l s t ab i l i t y of adding the caa rd su r f ace at Oo deflection at low st?ld moderate angles of aktack. For angles of attack greater t h m 15O et M = 0.60, the d i r ec t iona l s t ab i l i t y was s1ighti.y higher with the canard surface on. Deflection of the cemrd surface t o loo had l i t t l e effect 02 d i rec t iona l s t ab i l i t y izp t o m. angle of a t tack of Bo; however, above this alzgle of ettack, significant increases in d i rec t iona l s tab i l i ty occur red when the deflection was increased from Oo t o loo. (See f ig . 17.) Addi-Lion of the canard SUT- face caused large negative increnerrbs i n t o occur at moderate and

hlgh engles of' Ettack wi th the raxi$m effect occwring a t l5O. The r e su l t s of f igure 17 icdicate furt'iermore that these ef fec ts due t o addi- t i o n of the cm-z;rd surface were associated with interact ion on the wing rather than on the ver t ica l t a i l inesnuch es negative increments i n C

would not be q e c t e d t o eccompany gositive increments of Cyyp vhich came IYom E, ver t i ca l t a i l located above the rol l reference axis .

c z P

2P

The teil-or '? results presented in figures 11 2nd 12 show tmt some of the beneficial effects of deflecting the canard surface a t mderete

10 KACA Rl4 ~ 5 7 ~ 0 8

angles of a t tack can be a t t r i bu ted t o E small reduct ion in the ins tab i l i ty of t'ne wirg-body-cenard configuration. Zven wi th the cenard surface uniie- f lected, a small redxct ion in instabi l i ty with increasing angle of a t tack occurred w i t h the delta-wing configuration, and the trapezoid-wing model I

became directionally stable withoLt t i e t a i l a t high angles 02 a t tack ( f ig . 12). These ta i l -off ckaracter is t ics do zot follov the trends enccuntered on mny swept-wing"bo6y configarations whic3 generally show szbstatial i x r e a s e s i n d i r e c t i o n a l i n s t e b i l i t y a t noderately hi& angles of a t t ack? (For example, see refs. 5 an6 6.) The ta i l -off d i rect ional s tab i l i ty ckarac te r i s t ics of t'ne preser-t models are Sel ieved to be, appre- cFably influenced by two favorable effects not generally present on con- ventiozal mrangeLr.ents. O m of these cmfigurat ion effects i s the absence cf an efter'bo&y behind the wing imsnuch as re fereme 6 shows 2. favor&ble e f fec t of remving t k e afterbody- of a swept-wing-body arrangement. The other favorable effect i s believed t o occur a t the nose of the nodel and is mni les ted by a "strake effect" of %ne canard surface. The strake efTect w a s Tound in reference 7 from s r ~ l l horizontal strakes on a body nose xhic'n a l te red the forebo- flow s-dfficier"bly t o make a highly unstable s7cept-wing-boe arrangezerit becone d i rec t ioml ly s tzb le a t nigh angles. The sua11 favorable effect of cans& def lect ior , for the tail-off configu- ra t ion (figs. 11 ar-6 12, 1.1 = 0.60) cppears t o be furtker manifestation of the s t rake effect as indica%e& by the occ'zrence of posi t ive incremnts ir, both C: and CyB with posit ive increrrmts in control deflection. n3

The largest favorable effect of camrd Zeflectioc was on the t a i l coctr ibut ion to direct ional stability vhich i s s m i r l z e d i n figure 18. T'Aeee resu l t s show that @ins Ettribrtxble t o canard deflection were realized for both wing p k n f o r m and t k a t tie gains were such Kore pro- nounced f o r the Celt&-ving model vkic'n had a greater t a i l contribukion a t hi& angles of a t tack w i t h the canard swface deflected 10' t'nm at zero angle of attack. O n t he ot'ier :?and, -the t a i l coctribution t o dcrec- t i o m l s t a b i l i t y w i t h t'ne trapezoid wing decreased ragidly v i t h mgle of a t tack and becene negative 5-5 high angles. Reasons lor the differences in var ia t ions 03 tail contrLbutions with angle of a t tack For the two wizg plan form have not been establis3ed; however, it i s possible t h a t re la t ive ly minor differences in gemefry, such a s wing root leadiFS-edge location oc the body or exposed root chord length, had an FmportarrL ef fec t on tile flow a t the ve r t i ca l ta i l .

The exper'irnextal directiozai stabil i- ly resul-bs of the present study a t high subsonic speeds me in general agreeaent w i t ' s lox-speed r e su l t s obtained i n the p a s t i n f o r c e t e s t s i n the Langley rfree-flight t m z e l . (For e x w l e , see re3. 8 which presents resul ts la- a canard model having a 60' del ta wing. ) The resGlts of reference 8 Ciemonstrste t'mt &ding a camrd smface deflected 15' has a favorable efiect om the tail-off conTiguration and also increases the ver t ical- ta i l coztr ibut ion to the d i r e c t i m a l s t a b i l i t y a t low an6 moderate angles of attack.

NACA RM ~57508 - 11

Effect of cmerd s u f a c e and camr& deflection on aeroQnmic character is t ics in s idesl ip . - AerodynanLc character is t ics ol" the delta- wing model w i t h aDd without the canard surface are giver? i n figme 13 for angles or" at tack of anproximately Oo ecd and for a Yach nuioer of 0.60. The only s i g n i f i c a t e f f e c t s of the canard surface or canard deflection for an mgle of a t tack of approxiuately 0' was a nose-ap pitchilzg Eoaent md an increese in e f fec t ive diheEral which accompanied e. canzrd del"1ection of loo. At an angle of sttack approxinately 12.3' , an z-ppreciable nonlifiear variatiorz or" rolling-norrent coeflicient wi th sideslip angle occurred when the carlard surface w e s off . Tits f a c t indicated tha t a region of negative effective aihedral existed between mgles of s ides l ip of 51°. Addition of the caaard caused the var ie t ion t o becone l inear up t o bo uld geve the la rge cega t ive incremnt in C

a l so shown In the derivatives of f igure 17. 2P

The var ia t ion 'of L i f t , drag, and ~itching-noment coefficients w i t h sideslip presented i n figure 13 shows l i t t l e change i n lift and drag; however, a t an mgle of a t tack of apnroximlately 12.5O, the configuration w i t h the carard surface deflected loa shared a moderate var ia t ion of pitching-rnoxe-n-t coefr'icicnt w i t h sideslip angle.

The resulks of an investigation et high subsonic speeds of the s tEt ic longi tudiml m d lateral s t ab i l i t y c3a rac t e r i s t i c s of t w o c w r d airplane configurations are summrized in the followirq observations:

1. The longi tudiml control character is t ics indicated tkt s t a l l i n ! of the camrd surface a t ht-gh Oeflection angles (loo t o 1-5O> occ-mred zt re la t ive ly lo7$ model angles of et3ack which could inpose appreciable l imitat ions 011 t he mutinml t r i m lFft coeff ic ient -zttaicable f o r a longi- tudinelly stable configuration.

2. Significant itnpovetxents i n control effectivecess and maximum value 03 t r i m l i f t coefficient were ob-lained Tram d d i t i o n of a body f l ap having a conics1 shape, located slightly behin& the canard on tkke bottom of the body.

3. Addition 02 the canard surface a t Oo iieflection had l i t t l e e f fec t 011 overa l l d i rec t iona l s tab i l i ty of the Cielta-wing model, whereas deflec- t i o n of' the canard surface t o 10' had a k g e favorable effect on direc- t i o n a l s t a b i l i t y a t 3igh angles of a t tack fo r both the trapezoid- and delts-wing configurations tested.

12 NACA RM ~ 3 7 ~ 0 8

4. T'nere xere mrked d i f fe rences in overe l l d i rec t iona l s t sb i l i ty of %he delta-wing and trzpezoid-wing models w i t h the canard surface deflected LOo. DFrect ioml s tabi l l ty of t5e delte-wing model was greater a t high angles c~f a t tack thas &t t i e Lowest apsles, whereas the complete - model w i t k t 3 e t r q e z o i j . wing becme directiozally unstable at e xoder- e te ly 3igh mgle in spite 0: $he a2arernentioRed bell-efits of canard CeTlection.

Lmgley Aeromutical Ldcoratory, Netlor?al A6visox-y Comittee for BeronaxLFcs,

sangley Field, Va., September 17, 1957.

NACA E4 L57J08

REFEZENCES

1. Jobnson, Joseph L., Jr.: A Study of the Use or" Various Hi&-h--Lift Devices 02 the Horizontal T a i l of a Cmsrd Airplane Model ss E Kerns of I x r e e s i n g t'ne Allowable Center-of-Gravity Travel. N4CA RM ~521a83, 1953.

2. Driver, Corcelius: Longitudinal smi Leteral S tab i l i t y end Corr-Lrol Chsrscter is t ics of Two C m r d Airplane Configurations a t Yich Nwbers of 1.41 end 2.01 . K4CA RV ~ 5 6 ~ 1 9 , 1957.

3. Gillis, Clsrence L . , Polhasus, Edward C . , and G r z y , Joseph L . , Jr . : C h a r t s f o r Determining Jet-Boundary Correctiocs for Conplete Models i n 7- by 10-Foot Closed. Rectangular Wind Tunnels. NACA lWB L-123, 1945. (Fornerly NACA ARR Ls31.)

4. Herriot, JoBn G. : Slockge Corrections for T1.lree-DFnensional-Flo~J Closed-Throat Wind T-mels, With Consideration of Vie Effect of Compressibili-ty. NACA Rep. 995, 1950. (Supersedes NACA RM ~ 7 ~ 2 8 . )

5 . Fev, Albert G., Jr.: Investigation 2% High Subsonic Speeds of the S t s t i c La te ra l and Direct ional Stebi l i ty and Tail-loeds Character- i s t i c s of s. Model Baviog e. Eighly Tapered Swept Wing of Aspect Retio 3 .snd Two Horizoatsl-Tail Positions. NACA RM L56E29, 1936.

6. Polhmus, Edwara C., and S g r e e m , Kenneth P.: Subsonic Wind-Tunnel Investigakion of the Effect of Fxe lage Af'te-rbody on Directional S t ab i l i t y of: Wing-Fuselage Combinetion6 a t High h g l e s of 4"t;taclc. NACA TN 3896, 1956.

7. Sleemr?, Willian C., Jr.: Investigctioa a t Bigh Subsonic Speeds or" the Effects of Various Borizontsl Fuselage Forebody Fins on the Direct ioml md Longi tudinal Stabi l i ty of a Cozglete Mofiel %vil.lg a 45' Sweptback Wir!!. K4CA RM L5&-25, 1957. Y

8. Bates, Willim R.: Low-Speed Static Lateral Stabil i ty CharacterLstics of a Csnard Model Eaving e 60' Triangular Wing and Iiorizoctal T a i l . XACA E4 LgJI2, 1949.

14

3ody : Xaximuc? Biazeter. i n . . . . . . . . . . . . . . . . . . . . . 3-50 Len&.h. i n . . . . . . . . . . . . . . . . . . . . . . . . . . 37.00 Bese area. sq in . . . . . . . . . . . . . . . . . . . . . . . 9.582 Tineness r e t i o . . . . . . . . . . . . . . . . . . . . . . . 19.57

Trapezoid wing: Bgalz. i n . . . . . . . . . . . . . . . . . . . . . . . . . . . 25.72 Chord at bw-Xing ic tersect ion. in . . . . . . . . . . . . . 13.25 Area. sq f5 . . . . . . . . . . . . . . . . . . . . . . . . . 1-53 Aspect racio . . . . . . . . . . . . . . . . . . . . . . . . 3 Tagerracio . . . . . . . . . . . . . . . . . . . . . . . . . 0.143 Yhic:kness r & t i c . . . . . . . . . . . . . . . . . . . . . . . 0. 04 hfezc geometric c k o r ~ . in . . . . . . . . . . . . . . . . . . . L0.184 Sweeg argLe of le.=&tzg ebge. deg . . . . . . . . . . . . . . 38.67 &-ee> &@e or" trailiY4 edge. deg . . . . . . . . . . . . . . -11.3 Leading-e&ge half-mgle. mrml t o L .E.. keg . . . . . . . . 5

5 n Lrailing-edge hlf-mg1e. n o r r l t o T.E., de@; . . . . . . . . . Delce -%ing:

Spar. i n . . . . . . . . . . . . . . . . . . . . . . . . . . . 22.56 Chord az bow-ving in5ersection. 2- . . . . . . . . . . . . . 16.51 Keen geometrlc c?-ord. in . . . . . . . . . . . . . . . . . . . 13 . 027 Area. sq f t . . . . . . . . . . . . . . . . . . . . . . . . . 1-53 Aspect rc t io . . . . . . . . . . . . . . . . . . . . . . . . . 2 3 1 Thickness ratio . . . . . . . . . . . . . . . . . . . . . . . 0.036 Ledirg-edge klf-engle. n o ~ r . 1 t o L.Z., deg . . . . . . . . 5 Trailing-edge half-angle. n o r m l t o T.E., deg . . . . . . . . 5

CEnarc? : Cree ( t o t a l t o body cenzer 1Lr.e). sa_ in . . . . . . . . . . . 25.35& Ares. eaosed (eack cazwa). sq i n . . . . . . . . . . . . . . 6.7b2 Spm. emcse3. in . . . . . . . . . . . . . . . . . . . . . . 2. 25 Mean geozetric c.1ord. i n . . . . . . . . . . . . . . . . . . . 3-33 Za%ic of t o t a l ctirard ares t o t c t a l wiog exea . . . . . . . . 0.115

?krtical tai l : Area. expose&. sq PC . . . . . . . . . . . . . . . . . . . . 0.279 S g m . exposed. i n . . . . . . . . . . . . . . . . . . . . . . 4.25 Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . 0.439 Sveep of lee3ing edge. <e@; . . . . . . . . . . . . . . . . . 70 Sectioll . . . . . . . . . . . . . . . . . . . . . . . . 3/1.6-i~. slab LeeCing-edge i-&lI".mgle. zormsl t o L.E., aeg . . . . . . . . 5

Ven'crel f ins: k e a . each fi?.. eqosed. sq f+, . . . . . . . . . . . . . . . 0.13 Spa?. emosed. i n . . . . . . . . . . . . . . . . . . . . . . 2.25 Aspect r e t i o . . . . . . . . . . . . . . . . . . . . . . . . 0.271 Sweep of leaCing edge. deg . . . . . . . . . . . . . . . . . 63 Sveep of t r e i l i n g edge. deg . . . . . . . . . . . . . . . . . -77.5 LeeCing-edge helr".acgle. n o r r z l t o L.E., CeE; . . . . . . . . 5 Trailing-eQe ?xilf.zngle. ?loma1 t o I .3.. deg . . . . . . . . 5

.

Body s ta t ion

0 * 297 .627 - 956

1.285 1.615 1 943

2.605 2 936 3 267 3 598 3 - 929 4.260 4 592

2 -275

4 = 923 5 0255 5 -581 5 920 6.252 6 - 583 18.648 37 -000

RaCius

.a24 865 -903 .940 .958 996

1.020

1-07? sectioz 1.75

0 .076 .156 -233 307

$378

509 573 .627 .682

k h 4 5

-732 .780

16

Y Lateral force \

NACA RM L57J08

L I f l

A i pJ Yowmg moment

i

i Z

Figure 1.- Body reference axes showing posit ive directions of forces, moments, snd angular deflections.

A NACA RM ~57508

Figure 2.- General a,rr.zngenent of the basic model configuration showing the delta-wing and trapezoid-wing plan forns tes ted. (All d b n - sions are in inches.)

t-2/45 A

260 =&dto spreoder rod

Body flap Center mwnthg bracket

Figure 3. - Details of the body nose flap and simulated canard sp l i t flap tested on with the delta wing.

F P

ul I? 4

the model 2 03

I

25

20

J5

10

-a

c1

8: - 5 4

0

25

%O

15

IO

5

0

-5

.30

.25

-20

.I5

.IO

b b

6 0

0

.30

-25

.20

.I 5

.I 0

.05

0

-4 72 0 .2 4 .6 .8 ID L2 L i f f coefficient ,C' Li f t coef f ic ient ,G

Figure 4,- Effect of the canard surface on the longitudinal characteristics of the delta-wing configura.tion.

.I 0

.05

0

N 0

+s s E a E I

c CI

-..I" -4 -2 0 2 4 .6 .8 IO I2

h z a

0

0

0

-05

"IO

45

-113

-25

4' -5 0 5 IO I5 20 25 30

Li f t coef f ic ient ,C, Angle of attack , Q , deg

Figure IC. - Concluded.

I 9

Y '

"

r I

-30

I

25

20

I5

I O

b 2 . 5

b ,

x 0 * 2 0 P

0

-4 72 0 .2 4 .6 .8 I ! ! 12 -4 -.2 0 .2 4 .6 .8 I ! ! 12 L i f t meffIclent,CL LiFt coefflcient,CL

Figure 5.- Effcct of canard deflection on the longitudinal characteristics of the delta-wing configuration.

25

-20

.I5

. 10

.I! 5

0 .IO

.05

0

I :LJ -4 -2 0 2 4 .6 .8 ID I2

L i f t wef f ic ienf ,CL

10

05

0

:05

-IO

45

120

-25 -5 0 5 IO 15 20 25 30

Angle of offack, 0 , deg

Figure 5. - Continued.

.J 5

.I 0

-05

0

I IO

.05

0

YO5

-.IO

7 2 0 -4 72 O .2 4 ,6 .8 JD L i f t coefficient,CL

.J 5

.I 0

-05

0

I

Figure 5. - Concluded..

e; d 03

-5 0 5 JO 15 20 Angle of attack, Q , deg

w N

25

20

I5

lo 'b 8: PI - 5

0

0

30

25

.20

I5

.IO

05

0 3 9

0 .20

.I 5

0 .IG

.05

0

-4 -2 0 .2 4 .6 .8 I ! ! L 2 L i f t coefficient,CL

-9 12 0 .2 4 .6 .8 I! L 2 Lift weff icient,CL

Figure 6.- Effect of canard deflection on the longitudinal characteristics of the trapezoid-wing 03 configuration.

\

I c I I *

Figure 6. - Continued.

:*y4 :2 0 2 4 .6 .8 ID 12 L i f l coeffment ,CL

-5 0 5 10 15 20 25 30 Angle of attack, (I, deg

." -4 72 0 .2 4 .6 .8 ID -5 0 5 10 I5 20

L i f t mefficient,CL Angle of attack, u , deg

. I .

L

-0 P b

.rC

r, r,

B a x a Y h 4 T

25

20

15

IO

5

0

0

0

30

25

-20

.I5

.IO

05

0

0

0

30

25

.m

.I 5

.I 0

0

-4 -2 0 2 4 .6 -8 I.. L? L I f f coefficlent,CL

tc VI 4

0 4 [x,

F ' @ m e 7.- Effect of body nose flap and canard del-ta-wing conf igurE

flap on the longitudinal characteristics of ltion. M = 0.60.

.I5

io

D5

0

IO

05

0

.05

"IO

-15

-113

'2% -.2 0 2 4 .6 .8 ID I2 L I f f coefficienf ,C,

.I 5

.I 0

.05

0

I O

05

0

705

-.IO

-f 5

432

-.a -5 0 5 IO 15 20 25 30 Angle of affack, 0 , deg ul I?

4

8 Figure 7.- Concluded. (Flagged syllibols indicate 6, = 1.0'. ) a3

b 4

I I

25

%O

15

lo b

c

P - 5

4 I I 1

F

-4' -.2 0 2 4 .6 .8 fD I2 L i f t mefficient,C'

-4 72 0 .2 4 .6 .8 10 l2 L i f t coefficienl,CL

Pigurc 8.- Effect of the ventral fins on the longitudinal characteristics of the delta-wing configuration.

I O

05

0

0

0

0

-05

-.IO

-I5

- 2 3 , -5 0 5 IO /5 20 25 30 Angle of attack, u , deg

I I

”

I b

I 1

"_ 1 b

25

20

15

IO

b

b

P . 5

0

0

0

30

25

20

.I5

.IO

8 F -.- 05

0 .30 c,

9, D c,

.25

9 ?

0 .a

.I 5

0 .IO

-05

0

-4 72 0 2 4 .6 .8 ID 12 L i f f coefficient ,CL

Figure 9.- Effect of >-inch extension of the fuselage af'terbody on the longitudinal characteristics of the trapezoid-wing configurakion. 6, = Oo.

I

w P

”

-dw -4 72 0 .2 4 -6 .8 IO 12 L i f t mefflcienf ,C,

0

0

-5 0 5 IO 15 20 25 .?Q Angle of attack, a, deg

Figure 9.- Concluded.

I I

I 1

I

01

1 1

01

0 c%

- 01

Q02

001

crib

-001

001

0

-001

-002

-.Om

“004

0 CY b

-Dl

-02

002

,001

crib 0

- OOf

m1

0

-003

-004

-5 0 5 10 15 20 25 30 Angle of ottock,u, deg

-5 0 5 10 I5 20 At?g/e of attack, u , deg

.- .”

I * ;

P E

Figure 10.- Effect of the canard surface and canard de:rlection on the lateral s tab i l i ty derivaAives of the del-ta-wing conSiguration.

34

or

mr

0

0

403

NACA Ri ~57~08

-5 0 5 10 15 20 -5 0 5 10 15 20 Angle of ottuck, a, deg Ang/e of ottock, (I, deg

Figure 10.- Concluded.

1 I I

-5 0 5 IO I5 20 25 30 Angle of attack, 0,deg

-5 0 5 IO 15 20 Angle of affack, a , deg

Figure 11.- Effect of the verbical ta i l on the lateral stability dcrlvatives of the delta-wing configuration with the ventral fins removed.

01

0

-01

-02

002

.00 /

0

-001

.001

0

~ 0 0 1

cia ,002

-.004

01

0

- O/ - 02

002

DO/ .. '9 a

- oo/

.oo/

0

-.oo/

-.m3

-004

0

-O/

-02

.002

%! 0

-DO/

.oo/

0

-. 001

-m

-003

-.m4

NACA RM L57J08

-5 0 5 lo 15 20 Angle of ottock, f , deg

-5 0 5 lo 15 20 Angle of ottock, Q , deg

Figure 11. - Concluded.

Dl

0

701

002

.001

-002

.m3

.m2

001

0

002

I b

-5 0 5 IO I5 20 25 3 ~~

-5 0 5 10 15 20 Angle O f Uf fUCk, u,dq Angle of oftuck, a , deg

Figure 12.- Etfcct of both the vertical tail and ventrd. f i n s on the lateral stability derivatives of the trapezoid-wing configuration.

NACA RM L57J08

Angle o f aftuck, a, deg Ang/e of attack a deg

Figure 12. - Concluded.

C n

Figure 13. - EPfcct of the canard surface and canard deflection on the aerodynamic characteristics in sideslip of the delta-wlng configuration. M = 0.60.

- 4 -2 0 2 4 6 8 IO I2 I4 -4 -2 0 2 4 6 8 IO 12 I4

Angle of stdesllp, /, deg Angle of sldesltp,/, deg

(a) a = 0’.

I I

-4 -2 0 2 4 6 8 io 12 14

Angle of sideslip, a, deg

-4 -2 0 2 4 6 8 f0 12 f4

Angle of sideslip, ,8, deg

(b) a = 12.2'.

Figure 13. - Concluded.

I , L

6A NACA RM ~ 5 7 ~ 0 8

8

6

.002

0

.04

.02

0

41

Figure 1k.- Variation of stat ic longi tudinal and d i r ec t iona l s t ab i l i t y parmeters, lnfninuzli drag coefficient, and ~lraxFnu?l? lir”t-dreg r a t i o w i t h Mach nmber.

42

20

'p lo

0

- R L

06

-04

02

0

M = 0.60 M =0.94

10

5

0

NACA RE: L57JO8

-. / 0 .l 2 .3 0 .I .2 3 r i m / f f t coefficient, C,

Figure 15.- Variation of longitudinal characteristics for trim condi- t ions of the complete rr;odels h v f n g the delta wing and the trapezoid wing.

.

.IO E static morgin for & = 0"

l.

91 4

0 0

8 ; A 15 OBE stotic morgin fw &=OO

.. Y

12 0 .2 4 B 4 10 12 -2 0 .2 4 .6 .8 10 I.2 L i f t coefficient, C, L I f f coeffklent, C,

Figure 16.- Pitching-moment characteristics of the models showing control effectiveness of the canard with two values of low-lift static margin. M = 0.60.

!a

P i5

I w c

44

A

0

-.oo/

0

NACA RM L57J08

lncremenf due to conord "- Increment due to /Oocanord

deflectton

-004

Angle of otfock, a, deg

Figure 17.- Increments i n l a t e r a l stability derivatives due t o the canard surface and canard deflectior, for delta-wing configuration.

L

A

n

.

D

.Om

002

031

0

7001

.003

002

.oo/

0

.

.oo/

0

,001

Angle o f affock,a, d !

Figure 18.- ETfect of canard deflection on the %ail contribution to direct ional stzbil i ty for the delta-ving and trapezoid-wing models and ef fec t ol" the ven t r a l f i n on the d i rec t ionz l s tab i l i ty Tor the delta-wing configuration. M = 0.60.