Document
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COPY 2
RM L551271: Langley Aeronauticzl Laboratory Langle Field, V2.
""".."-
ATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
WASHINGTON February 17,1956 TEJTSONIC A E X O D Y W I C A N D TRIM C H A R " 1 S T L C S OF . A blULTI-ENCIXE DELTA-'WING ALFPlAKE MODEL By John M. SwihartandWillard E. Foss, Jr.
An investigation of' tiizee modelsof a delta-wing airplane designed r c fo long-range subsonic cruise end a supersonic dash has been conducted i n t h e Lal?gley 16-foot transonic timnel. A two-engine version of the
airpiane was tested a t e s s e n t i a l l y z e r o 3 d f t and it was found t o have a
high transonicdrag rise. The t w o four-engine nlodels were desip-ed eccordfng t o area-nile concepts to have Setter performance characteris- t i c s t h a n the two-eggixe model. The four-enginedelta-wingairplane models were similzr i n a l l respects except that one had a plzne delta wing w i t h full-span elevogs and the other had a cazzbered delta wing w i t h * partial-span elevons. Ti?e Mach number range of the imrestigation was I1ron 0.70 to 1.06 aad the Reynolds number range w a s r'rom 8.3 x 106 to 13.3 X 106 base& on w i a g mean aerodynamicchord. The angle-of -attack b range for the four-engine models was varied fron about - 3 O t o a velue necesserjr t o cbtain a lift coefficient ofabout 0 . 3 . Nadel t r i m char- a c t e r i s t i c s were obtaiped OE the focr-engine versions by deflecting the elevons I"rom O3 t o -2O arA -4O. Both four-engine models were t e s t e d with T 0 - u sFngle-engir;e nacelles end two tvin-engine nzcelles.
(The results of the iwestigation indicated that the four-engine de&tai- -=i:rplane models had onsiderably lower transonic-drag-riseincrements thm the two-engine model0 u t did not echieve the l o w transonic-drag-rise
Y
increxent of a sL-xLlsr four-enginedelta-wingconfiguretionwithfiearly parabolicaxialdistribution or' cross-sectiozalarea. The higher d m g rise or" the roar-engine delta-wing sirplane Todels com9ared with the more idealized fo;sr-en@.ne configuration was a t t r i b u t e d t o a n a u x i l i a r y h o r i - zontal aerofijmamic surface in close proximity t o the wlngand t o other dissimilarities incluiiing u h g incidence o f 3 O , canopy, increase ir? wing thickcess ratio, landing-gear fairings, afterhdy shape, and d i f f e r e n t d i s t r i b z t i o n s of cross-sectLonel area aboveand below the wing-&ord plsne.
I / - Tie c d e r e c ding xodels had higher values of rraxim l i f t - d r a g r a t i o
t k n t'ke$lzne-wingmodelsthroughout the Mach number range investigeted.
l
Tie cambered-wing twill-engine nacelle confFguration had ti.,e highest value .I of Fiaxinrurn l i f t - d r a g r a t i o for k c h numbers below 0.92. For trimned f l l g k t u i t h a s t a t i c rmrgin 3 percent of the neaEaerodynamicchord a t a -I- 7 I T\L L, coefficient GI" 0.23 (neer cruising l i f t coefficient), the canbered- wing twin-engine nacelie configuration had h i g h e s t l i f t - d r a g r a t i o and nearly constant elevm deflection overtile Nach nmber range investigated.
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E a r l i e r t e s t s o f E two-erqinedelta-wingairplane model d e s i a e d for E. losq-range cruising flight and e sxTersonic dzsh indica+,ed a high zero-liftdrag'risetkrough the transcnic speedrange.Inaneffort t o reduce t'ne drag rise of t h i s two-ergine airplane model, & four-engine delta-wing configuxtion (xodel 1 of r e f . 1) was designedbased on %he ccmepts of refereme 2 t o have an axial distr-ibution of cross-sectional area slxilar t o t:mt o f a parabolic body of revolltion with a f'ineness
ratio of 9 . Data f r o n t h e f r e e - f l i g h t test of this configmationindi-
cated a cor_sidera-Dly lover zero-lift drag rise t k n t h e o r i g i n a l two- engizedelta-wingairplacexodel.Furtherevidencetosugport the area- d i s t r i b u t i o n c o m e p t s for node1 l of reference l was ojtained when tests of a bodyof revol-Jtion kaving tne same axial d i s t r i b u t i o n o f cross- s e c t i o n a l area yielded almcst t h e same value of drag rise (ref. 1).
.T lnese resalts led t o +,he d e s i s of a Tour-engine delta-wing airplane nodel w i t h ar- area distribution very similar t o that 03 model 1 of ref- erenze 1 but IncoqoraTing sone deviations necessitated by practical aircraft design.
The e a r l i e r i n v e s t i g a t i o n of the two-engineiielta-wing a-irpLane n,odel to determire the zero-lift drag rise was made in the Lmgley 16-fost transonic tunnel aEd i s re-portedhereicforconqarisonpurposes.
Two versions OT the redesigned four-engice delta-wing eirplene model have Seen investigated a t l i f t i n g c o n d i t i o n s i n t h e Lacgley16-root . L ~..-nsonic V D ramel. Tne i w e s t i g a t i o n of tine four-engine delta-wing air- plane mdeis v&s t o d e t e r x h e t h e rise i n millixm drag coefficient with hhchnuxker azd t o evaluate the t r i m characteristics of several wing and r.acelLe c=r-Pigurp,tions i n ?.he transonicspeedrange. Tae effects of m c e l l e s and variousaerodycamicsurfacesattachedtothedroppablestore ("las5 mcixted beneath %he f u s e h g e ) or, the transcnic rfse of minimmi drag coefffcient were also inrestlsated.
The two four-enginedelta-wingairplane mdels were similar i n all respects except t:mt one b2d a plane delks wing with full-span elevons ar-d tha other Imd E canbered delta ving with partial-spanelevons. A cz-r$ereb-wing=ode1 was used f o r +,he present investigation because of
the expectedlower valxes ofdrag at l i f t i n g c o n d i t i o n s (refs. 3 and h >
zhztni$t be a t t a i n e d ir, cmrparisonwiththeplane-wing model. 30th the fo&ecgire models were tested with four single-engine gacelles and t v o twin-engine nacelles, hereafter referred to as "split" nacelles and "Shmese" nacelles,respecti-Jely. The Sianeseracelles were considered advantageczs from a power-package assessnent even though the nodel with These nacelles had a less favorakle cross-sectional ares. distribution t b a h t k e rcodel witt sclit nacelles and, therefore, might be sabject t o higher vzlues sf drag rise.
* A l l confi&urations, the two-enginenodeland the loar-engiEe models, were tested a t an elevondeflection of Oo. The Mach n i l ~ e r range f o r the J two-enginenodel m s from 0.80 to 1.06 and the corresponding Reynolds rumber range based on the wing near? aerodynamicchord was from 12.3 x 106 to 13.3 X 1 0 6 . The Mach number range f o r the four-enginec-odels w a s T r m 0.70 t o 1.06 and the Reynoids rider rangebased on wing neanaerodyxamic
chord was fron 8.5 x 106 t o 9.3 x lo6. Additionaltes-is were a l s o made
et elevon deflections of -2O and -bo f o r n o s t of the four-engine model configurationsinorder to t5etem"ine the mdel trim characteristics. The m a l e of attack for the two-engine model was mintained very near t o 0 i n order t o maintainzero lift over the k c h nunberrange. In general, the angle of attack of the four-engine models was vzried from about - 3 O t o a velue necessary t o obtain e lift coefficient of about 0 . 3 .
SYMaOLS A cross-sectionzl area AIR a s p e c t r a t i o I B base area L b wing span drzg coefficient, balance-masurercent drag coeff nacelleinternel-forcecoefficient,
m(vo - VE) - ~ ~ ( 1 3 ~ -
'0) cos(a + $ 1
%S lift coefficient, L/q$ pitching-moment coefficier-t f o r zero l i f t KACA FM L55127b
-
* rxesn aero&yn&.?icchord external drag L' sngle between nacelle center lir,e and Fuselage reference line l i f t model length pitching nocent a'oo.xt, 0 . 3 5 ~ ' Mach n-xher mass flow pointrass-flowratio, pEvz/pov0
pressurecoefficient,Plocal - Po
9 0 s t a t i c gressure dynamic p r e s s w e Reynolds mx!!er xing erea velocity distance from wing leadicg edge d i s t a n c e t o r e a r ofnose nodel angle of a t t a c k Keasuredfrom fuselage reference line d e f l e c t i m m-gle of elevon, positive down mss &er,sity : Slope Farmeters Subscripts : s base E n a c e l l e e x i t s t a t i o n I i n t e r n a l 0 f r e e stream except i n
%
MOD= DESC9IPTTON Two-EngineModel A sketch of the two-engicedelta-wingairplar-e model is shown i n I"i_g"-e 1 and dirensions 'for the model are given i n tab le I. Tie nodel was constructed o l mgnesiu?!and m~hogany. . Tce wLng had 2, d e l t a glan form w i t h the l e d i n g edgeswept back 65O and NACA 65~004 airfoil sec- t i o n s p a r a l l e l t o t h e plane or" symmetry. The fuselzgeana the droppable store (hereafter rer'erred to as e "pod1') were desip-ed to separate on the perting lize Shawn iz figure 1. A photogrzph of the tvo-engine delta- wing airplar-exodel m u t e d i n the tunnel is shown as f i g u r e 2. The nacelles were set et an Engle of incideme of - 2 . 1 3 ' w i t h respect t o the wing .
Figure 3 shows the nacelle configuration with the central spike i n l e t znd table Ii gives the nacelle and nacelle central spike dimensions.
Four-Engine lbdels
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Fuselage and droppable pod.- A scherztic diagrzx showing nodel d e t a i l s is gresented in figme 4 snd additiofial details of the model
- georretry are given in table III. The fuselage-sod combination sh0-0-
i n figure 4 is Indented f o r a Mach nmher of 1.00 i n accordance with thearea-ruleconsiderations ES given il? rer'erence 2. The dropppzble NACA RM L55127b pod i s attached to the undersysz-face of the I'usehge and the pod aero- dynanic ccntrols consist GI' a canard, pod wing, and pod v e n t r a l f i n .
Tlane-wing mdel. - The plane-wisg xodel kad 6g0 sveep of t'ne wing
leading &ge, -loo sweep of tire wicg t r a i l i n g edge, 3 O of incidence,an z s p e c t r a t i o of 2.1, zgd incor3crated NACA 000h. 08-63 a i r f o i l s e c t i o n s .
Sclid mgnesiuz construction v i t h a n overlay of bonded mahogany over the nidckord sections m s e q l o y e d on the delta wing whichhad full-spzn elevons acd provisions for rrounting e i t h e r S l a m s e or s p l i t r a c e l l e s .
Cambered-whg nodel.- Yce canbered-k-ing zodel vas similar i n d l resgects incladicg construetion x t e r i a l t o t h e plane-wing model except ?or theleading-edgecmker (15 percent of the local semispan, see fig. 5 ) and the smaller elevon area ofthecankered wing. The leading edge of the wing : e s drooped 2.86 percent of tke local senispan to pro- vide a n e a r l y e l l i p t i c a l ssanwiseloading a t the design l i f t cceffi- c i e c t (0.22) and k c h nmber (l.kl4). The elevons on Cne canibered wing exterided to 79.7 percent of t h e wingsemispan. For a d d i t i o n a l d e t a i l s of the ca,&ered-wing desip-, see figwe 5 and table I V .
STlit mcelles.- Fo-s separate nacelles (designated herein as "split 2acelles" t o decote EL single-ecgine d x t i n g systern) were Dounted i n an i d e n t i c a l mn7er on boththeplaneorcavbered wing. Tne s p l i t - n7acelle c o n f i g c a t i o n( f i g . 4) consisted of: ( a ) two inboard pylon- rmr;-n,ted r_zcelies suspended a t 40.30 percent of the wing semispan with t h e longitudiral axis p a r z l l e l t o t h e wixg chord (fig. 6 and table V), a.xd (b) two vjtboard nacelles nounted flush on theuppersurfaceof the Xing E t 64.63 percent of the sexdspar?and at an angle of incidence to t;?e wing chmdof - 3 O (fig. 7 and table V I ) . All s p l i t n a c e l l e s were mde of maszesim and had the same izternal contoar (table V >.
Sisxese nacelles. - Tke Siaqese nacelles (the t e r m "Siarese nzcelle"
denotes E rated pair of engine duc%iir,g s y s t e m ) were suspended from the lower surface of each wing parallel t o t h e chord plane on a pylon a t
45.72 9ercentof the semispan (fig. 8 and t & l e V I 1 1 . The same internal
ccntours were rmintained for the Siayese nacelle ducting as for the s p l i t nacelle.
Photogragks showing three-quarter' f r o n t views of t h e plane-wing xodei witi., s p l i t o a c e l l e s and t h e cmbered-wing model with Siamese nacelies sre given i n figure 9. Froct vLews of t h e sarx? configurations are presented i n f l a m e 10. Photographs showing the pod-wing location ecd the manner i n which the sod xing was f z i r e d t o t h e r a i n wing are given in figu-e 11. T5e t i p s e c t i o n s of t'ne f a i r e d pod wing are located i m i d e ",he landing-gear fairicgs as indicated by the dashed l i n e s on t h e ; ' i I -=-e.
TESTS Tests of a tvo-enginedelta-wi-ng aLr-plalzemodel and two four-engine delta-wi_ng airplane modelshave beenconducted i n the Langley16-foot transonict-mnel. The operational and flow c h a r a c t e r i s t i c s of the wind tunnel are given i n reference 5.
The force tests for the four-engine models were cocducted a t h c h
numbers from 0.70 t o 1.06 and et 3 e p o l d s numbers fro= 8.5 to 9.3 x lo6
(fig. 12). For t'ne two-engine mdel the angle o f a t t a c k was keptvery close t o 0 ' t o maintain zero lil't over a Yich n&errangefro= 0.80 t o 1.06 3rd a t Reynolds numers from 12.3 t o 13.3 x 106. In general, the sngle of a t t a c k of thefour-engine model w a s varied, at a given Mach number, from &out - 3 O t o e . value necessa-ry t o produce a lift coefficient of about 0 . 3 . This l i f t coefficient 'as sligitly higherthanthedesign transoniccmise lift coefficientof 0.25. The fourbasicconfiguretions (plane-wing s p l i t n a c e l l e s , plane-wing Siamese nacelles, cEzbered-wing s p l i t nacelles, and cankered-ving Siamese nacelles) were tested through- out t h e k c h number range for enelevondeflection of 0'. The plane- wing s p l i t - m c e l l e , cambered-wing split-necelle, and the cavhered-wing Simese nacelle configurations were s l s o t e s t e d st an elevon deflectior.
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of -2' and -4O. Aaditional tests were conductedthroughoct the Mach nun- ber r m g e at en elevon setting of 0 ' for the plane wing with oukboard
- s p l i t c a c e l l e s renoved and f o r the p l m e and cvnbered wings without
nacelles. There wzs a i r flow through the ducts for a l l nacelle-on tests on both t h e two-engineand thefour-engine models. A drag breakdown f o r the four-engine models wzs obtained 5y t e s t i n g the plane-wing s p l i t - nzcelle configuration with various pod conrponents removed (canard, pod
uing, andpod v e n t r a l f i n s h m i n f i g . I l ( a ) ). One test wes mde with
the pod wing faired t o the main wing f o r the plana-wing split-nacelle configuration (fig. l l ( b ) ).
Pressure tests t o determine base pressure Coefficient and nacelle interntil-force coefficient were made s h u l t a n e c u s l y w i t h the f o r c e t e s t s f o r the two-engiDe airplane mdel. Pressuretests were radeseparately Trom the force tests but throughout the sm-e angle-of-attack and Xsch n ~ ~ ~ k ~ e r range for the four-er?gine models with elevon deflections of 0 ' and -4'.
' Tae force data for the m d e l s t e s t e d were obteined from en i n t e r - rially rrrouq-ted six-conpcnextstrain-gegebalance.Fuselagebase-pressure I forces,nacellebase-pressureforces, and nacellein+,ernalforces yere detem-ined from pressllrerreasurenents.Internalpressures were neasured near t'ne e x i t s of b o t h m c e l l e s 02 the two-engineziodel and near the e x i t of me in3oard s p l i t r a c e l l e and one duct of e Sianese nacelle on t h e four-engine model. Choked f l o w was obteinedinthenacellesofthe four-engine models at Yach n u k e r s above 0.95 by using a r a t i o o f e x i t area t o i n l e t %rea of 1.13. T h i s ratioincoxbinationwith t h e exLernal nacellecontour resilltea in an anniiar base on eachnacelle. (See f i g . 6 . ) Data Reduction -An a a t o m t i c punch-cardsystem was used t o reduce the force data t o coefficien% ~ O E L All force data presented i n t h i s r e s o r t have been adjusted f o r base-pressureforcesendnacelleintermlforces.Figu-e 13 shows the n r i a t i o r , of pod base-gressure coefficient and necelle internal- force coefficient with Mach cur!ber f o r t h e two-enginedelta-wing airplane model.
The v a h e s of base-force coefficient, internal-force coefficient, ar.6 point r2ss-flow ratio for elevon settings of Oo are presented in f i g u r e 14 f o r the four-engine xiodels as a fw-ction of angle of attack for the test Mach n u 5 e r s . IE general, the base and i n t e r n a l f o r c e s were a5oul; the s m e a+, ~n elevon deflection of -bo as a t the s e t t i n g of 0'. The neasuredvalses of base-forcecoefficient and internal-force coefficient were used t o a d j u s t the force data and the - 2 ' elevon data were obtained from anaverage of the Oo and -4' elevon data.
N o correction :has been xzde f o r s t i n g tares. A t thepresent time, it is believed such effects would be small.
The angle of a t t a c k f o r the four-enginedelta-wingairplane model hes k e n corrected for balance an6 sting deflections and f o r a t u m e l upflow a s g u l a r i t y of 0 . 4 ~ tkt was i n v a r i a n t v i t h l a c k number.
I r . addition to the corrections already isdicated, the values of drag Coefficient presented in this report, other t h n the b a s i c deta (shown i n figs. 15 ';o 19) have been adjusted for tunnel-wall reflected- mve distwbances. This adjustxent was m%de by crossplotting the beslc drag data agdms', I k c 3 nmber and t h e n f a i r i n g a ciirve lower t k a n t h e basic data i n t h e h c h runherrangefrom 1.00 t o 1.06 where a m x i n m drag-coefficient-adjzstment value of 0.0015 a t a bel; n w b e r of 1.04 was employed. This val-de (O.OOl3) is based d i r e c t l y on e . correlation of f r e e - f l i g h t tests of the two-ergioe deita-wing airplane model and t h e present data for t h e two-enginedelta-wing airplane model a t zero lift.
r This sane adjustnent wzs assGTed t o apply for t'ne four-engine models and a t l i f t i n g c o n d i t i o n s .
I' m Accuracy The v a h e s p r e s e n t e d i n tiie followi??_g table indicete the maximum error that may be present, including a l l e r r o r s that were detected throughout the process of recording, reducing, and presenting the data ir_ i t s f i n a l form.
C L . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . fO.005
C j J . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +0.001
c , . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .20.0005
M . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ko.005
a , d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . *o. 1
rJ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . kO.005
The r e p e a t a b i l i t y of the aerodynamic coefficients w a s better than the indicatedaccuracy. It was found t b a t the dragcoefficient, f o r exewle, repeeted within kO.0005.
Tne resdts of the investigstion are presented i n figures 15 t o 31.
Basic aerodynarnic c k c a c t e r i s t i c s are presented f o r a l l of the four- enginedelta--wingai-rplaze model coofigwetions tested. The z e r o - l i f t drag data for the two-enginedelta-wing airplane nodel ere presented in- conqarison with t h e Tour-enginedelta-wing airplene-mdel drag data.
The m r i a t i o n of slope paraxeters w i t h Piicn nm3er is generally shown f o r only the fou-V-engine basicconfigurationstested. The lift-curve and pitching-nomnt-curve slopes were obtained fro= straight l i n e s averaging that gortion of the curvesbetween a l l f t coefficiefit of 0 and 0 . 3 . All reference t o the transonic rise i n d r a g c o e f f i c i e n t i n the following discussion i s f o r the &ch n M e r range from 0.9 t o 1.04.
The free-strezm-tube area contzining the mss flow entering t k ducts at M = 1.0 hasSeensubtracted fro111 a l l areadiagramspresented. An indication of the data presented i n figures 15 t o 31 is given i n the following table: Figure
Basicaerodynamic c h a r a c t e r i s t i c s (a, CD, and C , egainst CL) . . 15 t o 19
Drzg-riseplots and area diagrams . . . . . . . . . . . . . . . 20 to 22
Effect of wing ca&er 02 drag coefficient . . . . . . . . . . . 23
Effect of c a ~ e r on drag due t o l i f t zt M = 0.90 . . . . . . . . 24
VaFiation of m u r ~ m r r ! l i f t - d r a g r e t i o ana l i T t c o e f f i c i e n t for
r e x i . ? l i f t - d r a g r a t i o w i t h Yach number . . . . . . . . . . . 25
Variatlon of l i f t - c - m e slope w i t h Each nurllber . . . . . . . . . 26
BACA REI L35127b Figure V a r i a t i o l -with Kachnumber of the zero-lift pitcking-monent coefficient, the slope of the pitching-norent coefficient agsixst Lift-coefficient curves, and the pitching-mrent c o e f f i c i e n t a t a l i T t c o e f f l c i e r t of0.25 . . . . . . . . . 27 t o 29
Elevon e f f e c t i v e z e s s p a r m e t e r s . . . . . . . . . . . . . . . 30
T r i m c h a r e z t e r i s t i c s . . . . . . . . . . . . . . . . . . . . 31
DISCUSSION AerodymxicCharec-ceristics os" Models cf minix-m drag coefficiect with Mach nwher.- One of the Variation purposes of tae present investigation was t o determine the transonic drag-rise characteristics of a long-rangedelta-wing a i ~ l e n e nodel.
E z r l i a r tests of the original versicn of the airplane, the two-engine delta-wing airplane rrodel condxcted ill t h e Langley 16-foot transonic t m n e l ,i n d i c a t e d a highzero-liftdrag rise. Specifically,these tests of t h e two-enginedelta-wixg a i r p l a n e m d e l show a zero-lift transonic drag-rise of 0.C212 (fig.23(a) ). I n an e f f o r t t o reducethetransonic d r a g r i s e of t h i s a i r p l a n e nodel, an idealizedfour-enginedelta-wing configa-ation was designed t o have an wrial cross-sectional erea dis-tri- SutioE elnost t i e s8ne as that of a parabolic body ofrevol-Jtion heving a f i c e n e s sr a t i o of 9.0. T3is attempt a t tracsonic-drag-risereduction vas based on the results of reference 2 which showed t h a t t n e z e r o - l i f t drag rise of a wing-%ody coxbinatior, could be reduced by designing the confi,mation to have a gradual axial increese and d e c r e a s e i n t o t a l cross-sectional area andbykeepingthe maxim? cross-sectional area t o e m i n i m . Tne r e s u l t s or" t h e f r e e - f l i g h t test of the fou-r-engine delta- 1 of ref. 1) hdica-ted a zero-lift drag rise ring configuration (nodei of 0.0100. The v a l i d i t y of area-ruleconcepts ir, thedesign of model 1 of reference I was F x t h e r e s t a b l i s h e d by achieving the same drag-rise increxent (0.0100) with a Cody of revolution having an axial cross- sectionai area distribution ideEtica1 with tht of model 1. These results led t o t h e p r e s e n t desig.1of the four-engice delta-wing airplane ;node Is.
Thesefour-er,gir,e xodelshavean are& d i s t r i b u t i o l approaching ,that cr" mcdel 1 of reference 1 OP a total cross-sectional area basis. The nondimensioml area progressions of khese models are Fresented i n fig- ure 20(b).. The Cifference il nondixelsionslcross-sectional area between nodel 1 of reference 1 a r d the Fresent Tour-eggire delta-wing airplane rnodel w i t h s p l i t m e e l l e s i s p r L m r i l y t h e r e s u l t o f a difference in w- f i n e n e s s r a t i o Setween the models (9.0 and8.2, respectively ). Some difference is a l s o dze t o a more rearward location of the maximum cross- s e c t i o n a l area.
A transonic drag rise ol' 0.013h w a s meesured for the plane-wing t split-necelle model End 0.0175, f o r t i e same wiag with Siamese nacelles
as show- infigure20(a).Inspection of theareadiagrsms(fig.20(b 1)
i n d i c a t e s t h a t t h e s p l i t n a c e l l e s bave the more favorable area distribu- t i o n and, therefore,tliesenacellesshould 'nave a lowertransonfcdrag r i s e Chm th% Siamese nacelles. Although thefour-engine deltz.-wing eir- plane models show appreciableredwtionindrag rise when conpared v i t h t h e two-engine nodel, they have about 50 percent higher drag rise than =ode1 1 ol' reference 1. Tb-e reason for the feilure of' the present nodels t o achieve <?e low (0.0100) brag rcse increnent of d e l 1 canbe explained by noting the ghysical differences between t h e Eodelsand the r e s u l t s from thepresent drag-breakdown tests. The physicaldifferences between the models are as follows: Present Node1 1 model of ref. 1
Pod wing . . . . . . . . . . . . . . . . . . . . . Yes None
Incidence,deg . . . . . . . . . . . . . . . . . .
3 None FLneness retio (equivalent 'OW 1 . . . . . . . . .
8.2 9.0
Landing-gear f a i r i n g s . . . . . . . . . . . . . . Yes None
- Cznopy
. . . . . . . . . . . . . . . . . . . . . . Yes None
Ca-rd . . . . . . . . . . . . . . . . . . . . . . Yes None
Ying thic-hess, percent . . . . . . . . . . . . . 4.08 3.0
II
Afterbodyshage . . . . . . . S l i g h t l yd i f f e r e n t diameters andslopes
The four-engine rrodels differed also In their distribution of cross- sectionzl azee i n that tb-e present rrodels 3sd an urrsymnetrical d i s t r i b u - t i o c e5ove and below t5e wing-chord plane (Fig. 20(b) ), wherees model 1 had E nearly symnetrical dfstributioo.
The r e s u l t s of d-ag-breekdown t e s t s f o r t i e podcomponents 0 1 1 the
plane-wing sglit-racelle configuration (fig. 21(a) 1 igAicate t'iat t h e
_uod %ring is thechiefcontributor t o tb-e transonicdrag rise. 1% a6as
an increnent in the rise of mcnimm drag coefficient of 0.0020 and, i f based 0 1 1 its own aresl the drag rise of: the pod wing would have the unreasonaslyhighvalue or" 0.0206. Shadowgraphs taken during the tests h d i c a t e a skock fornation mar t h e t r a i l i n g edgeof the pod wing that is not present for tests withoutthe pod wing. Fairing tke pod wfng t o the main wing ( f i g . 2 1 ( a ) ) r e d x e d t h e rise irr minimum drag coefficient by about 0.0010. It can be seen in figure 21(b ) t h a t t h e removal of t h e pod wicg lowers the area diagram s l i g h t l y i n e region of high slope, but
-
the difference in drag from erea-rule coxx,iderations would be less t32n 0.0020. The 0.0020 increment Fn dragcoefficient emphasizes t h e poir"L w tb25, when aerodynm-ic conponents are added t o e configuration, high local interferences pay occw, seperetion m y also OCCUT, and f o r t h i s particular surface choking of the flow between the pod wtng and the main wing ray result i n added drag.
NACA RM ~ 5 x 2 ~ 0
-
If the increment in transonic drag rise due t o t h e pod wing (0.0020) i s s ~ c t r a c t e d from the drag rise of the coxplete configuration (0.0154 f o r ti.,e plane wing w i t c s p l i t n e c e l l e s >, t h e r e s u i t i n g t r a n s o n i c d r a g r i s e would be 0.3134 o r 0.0034 higher than t h a t f o r xodel i of reference 1.
A n a2alysLs made to evaluate the incrercnt i E drag rrse due t o each of t h e r e m i n i r g p h y s i c a l d i f f e r e n c e s between the present four-engine xodels and nodel 1 indicated that no large drag-rise increxent (none of the rag- nltude of the pod wing a t l e a s t ) could be emected for any one dissimi- larit;r and, in general, it vas found tiiat each i x r e x e n t was within the accuracyof t'r-e reference dats. Since the differences between t h e rnodels (other than the FCC -ring) are such that an increase in drag rise f o r t h e present fom-engine models woxLLd beexpected, it is believed that the higher drag rise (compared w i t h that of nodel 1) would be adequately e q l a i n e d i-f' each dissimilarity Increased tP& d r a g r i s e by as small an increment as O.OOO5.
In general, the nacelle drag increnent shown i n f i g u r e 2 2 ( a ) indi- cated EO adverse nacelle interference effects.
I n theprecedingdiscrzssion,thecoxparisons of the data from the four-engine delta-wing airplane xodels of the present test with that o f reference 1 have beer- confined t o theplane wing. P i e can3ered wing could have beenused for the comparisons a l s o s i n c e it shows alnost the sane r i s e i n minimum d r a g c o e f f i c i e n t f o r a giver! nacelle installatior- as theplane wing. (See f i g . 20(2.). ) Calculations of the zero-lift drag rise f o r s e v e r a l of the cqnfigu- rations teated using the xethod of reference 6 were consistently Lower t h a n the neasured results, often by as much as 50 percent. It i s believed tkt t h e i n a b i l i t y of the rethod to give accarate predictions i s r e l a t e d to the zbrupt chw-ges irs sloDe o f the area diagrams for the configurations being icvestigated, and t h e i n a b i l i t y oftne nEthod t o i n c l u d e t h e e f f e c t s of segaratedflow and choked flow. The xethodacczratelypredicted tk;e z e r o - l i f t d r a g rise of T-odel 1 of reference 1, which had a f a i r l y smooth cross-sectional area progression.
Variation of CD with &ch number a t CL = 0.25. - Model 1 of r e f -
erence 1 and the twc-engke delta-wing aiqlane model were not investi- gated at l i f t k g conditione;. Tke l i f t coefficier;t of 0.25 was chosen f o r t k drag-coefficierst data. ?resented in figure 20(a) because it repre- s e n t s t h e d e s i g l l i f t coefficient Tor transonic cruising flight of the four-enginedelta-wingairplaqe models. A t a l i f t c o e f f i c i e n t of 0.25, t h e e f f e c t of tlrle type of nzcelle or? the transonic-drag-rise increnent is the stme as that a t t h e rL.nirxm drag coefficient; that is, the Siamese nacelles on either piace cr canioered xing rreiritain an increrrent i n d r a g rise of abotlt 0.0020 over Kmt os" the s p i i t nacelles. An i n t e r e s t i n g aspect of the t r m s o n i c d r a g r i s e under l i f t i n g c o f i i t i o n s is the% the plane wing r i t h a given rstcelle :has a lover &rag r i s e t h a n t h e same * configuretion hes at mini;nun drag coefficient. The carbered wing with a given nacelle installation, however, has a higher dreg rise w d e r lifting conditions then that of the sane configuration at xLnimx &rag coefficient.
The absolute drag-coefficient level is lower for the cmhered-wing models at t'r?is lift coefficient than for the slane-wing models for the lhch nu"- ber rage investigated.
Effect of cax-her on drag coefficient.- The benefits of camber in reduclng the drag coefficient at lifting conditions are readily shorn- in figure 23. In this figme, the drag coefficient of tke ca7;bered wing 3as been sdotracted from the drag coefficient of the p U e wing at a given Yach nuxber for lift coefficieats of 0 . 1 5 , 0.20, and 0.25 m d the results plotted sgainst Mach nm3er. 1% may be noted that, subsonically, c-er provides a reduction In drag coefficient of about 0 . 0 0 4 0 for eitlner nacelle installation at a lift coefficient of 0.25. As would be expected, the beneficial effect of cam5er on the drag coefficient decreeses with decrezsing lift coefficient. At a Ylch l l u r d b e r of 1-04, %he ca?lbered wing still has z m advantage over the plene wing but thcs advantage is reduced at all lift coefficients.
Effect of c d e r on the drag due to lift at a hkch number or' 0 . 9 0 . -
The beneficial eTfect of czzzker on the &rag coefficient at selected lift
-
coefficients b s been shown in figure 23 over the Yach number range. It is of interest to show the effect of camber on the drag due to lift et 2 . Mach mmber of 0.90, the selected subsonic cruise speed.. for this desip-.
Sho-m in figure 24 is the dreg coefficient plotted against lift coeffi- cient for the plane a-n-d cm3ered wings with no cecelles, s p l i t nacelles, m8 Sienese nacelles. For comparison purposesy e . curve representing the
minim? possible induced drag coefficient ( X . = c . ' ) is shown passing
rim through the poifit for zero-lift drag for tie plane wing. It is assuned that the zero-lift drag coefficient for the plane wizg reyesents the skin-friction drag for the configuration end tl.at the additional drag coefficient shorn- for the caxibered wing Et zero lift is the increment in drag due to twist and c d e r . These data show thzt the caribered wing without nacelles bas almost the minimum possible value of induced drag coefficient. The data of reference 1 : for a caniberea delta w h g of aspect ratio 2 indicates this sane result up to tbe design lift coefficient of the -Jfag.
There is an increase of about 22 percent over the minimum possi- ble vzlue or" induced drag coefficient for both mcelle configura- tions at e . lift coefficient of 0.25. This increase in the drag due to lift is probably the result of the nacelles and fuselage ceusing w distortcon of the design elliptical spa^ load distribution on the wing. Reference 7 indicates that there should be no iiecrease in the drag due to lift when the test Reynolds number is 10 x 10 6 or greater. If the plane-wing zero-lift drag coefficient (0.0133) is
1k t XACA RM ~ 5 - 5 1 2 ~
comerted to the skin-friction drag coefficie2t by t h e r a t i o of model wing %rea t o wetted area, the skin-friction drag coefficient is 0.0033.
Tie v a P x of the skir-friction drag coefficient for the flat p l a t e at
M = 0.90 and R = 10 x lo6 is 0.0028 (ref. & 1 . This low value of skin-
friction drag coefficient thus precludes arq large interferences thet Ir:igi?t ca%e segarationdrag.Sincethexodelwith no nacelles has almost the nicirnm gossible value of drag due t o l i f t , it would not be e q e c t e d t%t an iccrease in Reynolds nmber would result i n a decrease in the drag dae t o lift.
Vzriatior. of (L/D and C L ( , . / ~ )ym with &ch number. - The
__L oariztior, of (L/D)nax with Kkch number (fig. 25 1 shows that the cam- bered wing with e i t h e r s p l i t or Siaxese nacelle has a higher value of ( L / D ) , , thsntheplane wing witheithernacelle. Below a k c h number of a3out 0.92, t j e Siaxese nacelles on either wing ’nave a s l i g h t l y higher (L/D)rax thar? t h es P l i tn a c e l l e s on e i t h e r wing. The highest vzlue of (L/D (that is, 10.7) for tf;e principal configurations tesked occurred for the cambered-wing Siarese nacelle configuration a t a Mach number or” 0.70. For &ch nu?lbers above 0.97, t h e cankered-wing bas hig5ervalues of (L/D)- thanthe splif-zacelleconfiguration cambered wirg with Siaxesenacelles. Below a Mach number of 0.95, t h e cambered-wing nacelle corfigurstions attain (L/D)- a t a CL of about 0.24 (fig. 251, Thereastheplane-wingnacelleconfigurations
att;eir_ their (L/D >rrax a t a CL of about 0.20. The added advantage of
thecankered wing is also realized i n notingthat i t s ( L / D ) , , . = occur6 a t nearly -the c n i s i n g l i f t coefficient o f t h e aircraft.
VariaTioEof C h with Mach nunher.- In general, the variation of thelif%-curveslope C L ~ withFfch nzmber is a3outthe sane f o r the fourprincipalfour-engineccnfigurations tested (fig. 26). W.e l i f t - curve slope C h varies fron abollc 0.045 subsonicallytoabout 0.037 near Mach r-uber 1.0. Far Mach mmbers up t o 0.96, t h e sglit nacelles have a slightlykigkervalueof C b t l a n t h e Siamese nacelles OJ? e i t h e r wing.
Variation of ‘JQ, &,/&L, and CwL=o.25 with Mach number.- The
I zero-iif% 3itchilg-moment coefficient generally shows a mall nega- tive increase with Mach nnnber up t o M = 0.95 (fig. 27). For Mach nun- bers higher t&m 0.95 and - ~ p 50 M = 1.0, the split nacelles on e i t h e r
wing producehighernegativevalues of 12%. The Sianesenacelles’ on
e l t h e r wing, bo-ever, h&ve lcwernegativevzlues of CTQ f o r Mach num- bers above 0.95. T?e differenttrend ir t h e v a r i a t i o n of Cro w i t h Mach I nunioer f o r the *do types of nacelles issholm in a trirn analysis in 8 scbsequent section of this pEper to have an ixportant bearing on nacelle a selection.
Tine variation of & , / & ' L with Ikch m&er (fig. 28) indicates that the foilr principal four-engine configurt=tions tested ere longitudinally imstzble &tout the 35-percent rrean aerodymnic chord up toa Mach nwber of atout 0.95. Beyond this hkch nunher, the configurations becone stable, with the Siur.ese nacelle configuration iadicating the greatest degree of stability above M = 0.98. Both the split and Sianese nacelles produce a clestz'cilizing effect since the configurations without nacelles ere stable at practically all k c h numbers.
The gitching-moment coefficient a t S = Oo (fig. 29) is approxi- mately zero for h c h numbers below 0 . 8 0 end increases negatively from M = 0 . 8 0 to M = 1.04 for a l l fo-=-engine collr"igurations. Elevon deflections of - 2 ' end -4* increase the pitching mxent at all &ch nun- bers for all coD3igurations. At M = 0.90, approxilritely -lo of elevon deflectior, would be required to trim about the 0.35~' for the split- nacelle configurations.
Effect of Elevoo Deflection on A e r o d m c Characteristics o f ' t i l e F o u r - & @ n e Model Variation of CQ and C , * with hkch number.- In figure 30 the plece-wing split-nacelle configuration shows higher lift effectiveness than the cadered-will-g split-nacelle cosfiguration. Tnis increase in lift effectiveness parmeter is zttributed to the larger elevon area on the plane wing. 'The cC!oered-%5ng Sianese nacelle configuration 'nes the lowest value or" lift effectiveness paraneter up to a Mach cumber of 0.95 but the parameter does not decrease for this coIll'iguration until e Mach rxnber of 0.97, vhereas t ' n e p " e - and caxiiered-wing models w5th split naceiles show a decrease in t'ne paraneter above 0.95. Tiis earlier loss in lift effectiveness for the models with s p l i t nacelles may be associ- ated with shocks neer the bsse of the outboard split nacelle.
The -plane-wing split-nacelle configuration has the largest value of pitching-moment effectiveness parameter up t o a Mach number or" 0.975.
The pitching-noment effectiveness paraneter follows the same trends with k c h nurioer as does the lift effectiveness permeter, the cdered-wing Simese necelle configuration hving the largest values above a Ekch murber of 0.975. Theoretical considerations would indicate that the mxlmuz value of these parmeters would occ-ur at a Mach nuxber of 1 . 0 bxt the data indicate that, for configurztions where shocks may be loceted on part of the elevons, the effectiveness of the elevons w i l l decrease at Vach nuaers of less than 1 . 0 .
16 r NACA TZM ~ 5 5 1 2 ~
.
ETfectof Mech number on elevon angle, angle of attack, drag coer'fi- cier?t, and l i f t - d r a g r e t i o a t t r i m . - The trbr a m l y s i s p r e s e n t e d i n f i g - P ure 31 for three of %he four-engine configurations tested is based on a constan+,3-percent-c' s t a t i c m r g i n ar,d a l i f t c o e f f i c i e n t of 0.25.
A t Elach nwbers above 0.90, the elevon mgle required to trim the caTbered-andplane-wing xodels w i t h s p l i t n a c e l l e s changed rapidly until anelevonangle of about -3.5' w a s needed a t a Yach number of 1.00. The canbered-wingSiamese-nacelleconfiguration, however, required a rela- t i v e l y srall change i n elevon angle for t r i m over the Mach nwfber range.
It was noted previously that t h e z e r o - l i f t p i t c h i n g moment f o r the Sienese ricelle configurations Secame less negative et Mach numbers above 0.90.
This decreasein ( 2 % requires a smaller elevondeflectionto t r i m t h e cambered-wing SiaTese nacelle coofiguration.
The angle-of-attack variation with Yach nur5er a t t r i m is about 1 ' for a l l configurations. The c&Tkered-wing models, however, require a higher angle of a t t a c k t o t r i m than the plane-wing models.
It was noted i n %he discussion of the drag coefficient a t a l i f t coefficier,t of 0.25 that the drag rise was 0.002 g r e a t e r for the Si&., wese nacellesthanforthe split nacelles. A t trim, however, the cambered-wing Siamese nacelle configuration hes the lowest value of drag coefficient at all Mach nurnbers beceuse, as indicated previously, this configuration requires less elevon deflection for tri,med f l i g h t . The drag-coefficient I data presented in figure 31 were obtained by crossplotting the t e s t data.
The l i l t - d r a g r a t i o at t r i m f o r t h e canbered-wing model with Siamese the lower drag coefficient for this configaration and n a c e l l e s r e f l e c t s is generally higher over the k c h m d e r range than that of the other rnodels .
CONCLUSIOMS An investigation of a two-engize delta-wing airplane model with no and of two four-engine delta-wing airplane models fuselage indentation (the design of which w a s based on area-rule concepts) in t h e Langley 16-foot transonic.tunne1 has indicated the folloxing conclusions: 1. The four-enginedelta-wingairplaneEodelswithfavorable axial d i s t r i b u t i o n s of cross-sectional area had considerably lower transonic drag-rise increments than the two-engine delts-wing airplane model.
2. The four-enginedelta-wingairplane models had cross-sectional &rea d i s t r i b u t i o n s sirdlar t o an idealized four-engine delta-wing con- f i g u r a t i o n with a nearly parabolic distribution of cross-sectional area .
but did not lzsve the lox ninimun drag rise of the latter beczuse 03 an auxiliary pod surface in close proximity to the wingand other dissimi- a l a r i t i e s i n c l u d i n g wing incidence of 3 O , callopy, increese in wing thick- cess,Iznding-gearfairings,af'terbody sllape,and a r e a d i s t r i b u t i o n s sbove er-d belox the wing-chord plane.
3. The cadered-wing xodels hzd higher values of maxim l i f t - d r z g r a t i o than theplme-wing models throughout the Mach number range.For Mzch numbers up to 0.92, thecakered-wingSianesenacelleconffiguration had s l i g h t l y higher values of rexinun l i f t - d r a g r a t i o tbn the cambered- wing split-Faceile conrigmation.
4. Tie carbered-wing models hed lower dreg coefficients t'om the et l i f t coefficients from 0.15 t o 0.25 throughout the plane-wing m d e l s test hkch number range.
5 . The c a ~ e r e d - w i n gmodel with 30 nacelles has nearly the m i n b r possible velue of' induced drag coefficient at a Ekch number of 0.90 and there i s only a srill increase in the drag due t o lift when the nacelles a r e added. The data indicate that there w i l l be EO reduction in drag due t o l i f t when the Reynolds n-mrer is increesed above the test velue of zpproxinately 10 x 16.
-
6 . T o r trirrmed ?light u i t h a 3-percent s"etic lnargin and a lift % coefficient of 0.25 (cear tk.e desi- cruising l i f t coefficient 1 , the cambered-wing Siamese ozcelle cosTigxation had highest values of l i f t - &rag r a t i o end nearly constant elevon deflection throughout the Ehch number rangetested.
Langley Aeromutical Ieboratory, Iiztioral Advisory C o m i t t e e f o r Aeronautics, Langley Field, V e . , Septenher 15, 1955.
XACA RM ~ 5 5 1 2 ~ c REFTWRCES C 1. Hopko, R;lssell N . , Piland,Robert O., and kll, Jazes R.: Drag Meas- a r e s e n t s a t Lo-+T L i f t of a Four-NacelleAirpianeConfiguration Having a Longitudinal D'rstribution of Cross-Sectional Area Conducive t o Low Transonic Drag Rise. NACA RM L53329, 1953.
2. 'hitconb,Richard T. : A Study of theZero-LiftDrag-RiseCharacter- i s t i c s of WLng-Eody Combinetions Fear t h e Speed of Soulld. NACA RM L52HCl8, 1952.
3. Jones, Robert T.: Estixated Lift-DragPatios a t Supersonic Speed.
NACA TN 1350, 19h7.
4. Hall, Cllarles F.: L i f t , Drag, and Pitching Mment of Low-Aspect-Ratio Wings at SJbsonic and SugersonicSpeeds. NACA RTvl A53A30, 1953.
5. Ward, Vernon G., TMhitco%b, Charles F., andPearson, Merwin D.: Air- Flow and Pover Characterfstics of the Langley 16-~oot Transonic Tmcel With S l o t t e d Test Section. EACA REI L52E01, 1952.
6. Nelson,Robert L., aEd Stoney, Willisx E., Jr. : Pressure Drag or' Bodies a t h c h Numbers UI, t o 2.0. K4CA RM L53I22c, 1953.
7. CsSorce,3obert S., andKelly, Thomas C. : A Note on t'ne Drag Due t o L i f t of Delta Wings a t Mach Nunhers ap t o 2.0. NACA R M L53Al6a, 8. Van Driest, E . R.: Turbulent Eoundary Layer i n Corrrpressible Fluids.
Jour. Aero. Sci., vol. 18, no. 3, M a r . 195l, pp. 145-160, 216.
t NACA I 3 1 L55127b .
a Whg :
Area. sa_ i n . . . . . . . . . . . . . . . . . . . . . . . . . L . 728
Span. i n . . . . . . . . . . . . . . . . . . .
. . . . . . . 56.721
Root chord. i n . . . . . . . . . . . . . . . . . . . . . . . 60.874
Length of M.A.C., i n . . . . . . . . . . . . . . . . . . . . 40.583
A i r f o i l s e c t i o n ( p a r a l l e l t o plane of symmetry) . . . . . X4CA 65A004
Sweepback lezding edge.deg . . . . . . . . . .
. . . . . . . 65
Dihedrel . . . . . . . . . . . . . . . . . . .
. . . . . . . -20 27f
Incidence, deg . . . . . . . . . . . . . . . .
. . . . . . . 0
Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . 1 . 8 6
FuEelage :
Overall length. in . . . . . . . . . . . . . . . . . . . . . 80.00
Distance from nose or" fuselage t o leading edge of wing
r o o t chord. i n . . . . . . . . . . . . . . . . . . . . . . 12.825
Veximn width. i n . . . . . . . . . . . . . . . . . . . . . . 5.980
Pod:
Overall length. in . . . . . . . . . . . . . . . . . . . . . 90.162
Distance from pod Ease t o leading edge of wing r o o t
c h o r d . i n . . . . . . . . . . . . . . . . . . . . . . . . . 21.025
lrbxirrmm width. i-r? . . . . . . . . . . . . . . . . . . . . . . 6.000
Waceiles :
Oirerall length. ir- . . . . . . . . . . . . . . . . . . . . . . 42.621
Bit d i a t e r . i n
. . . . . . . . . . . . . . . . . . . . . . 3.252
D i s t m c e from a l r p k n e cen_ter l i n e t o a a c e l l e c e n t e r
l i n e . i n . . . . . . . . . . . . . . . '. . . . . . . . . . 16.350
Vertical tzil:
Total wee. sq i n . . . . . . . . . . . . . . . . . . . . . . 125.271
Span. i n . . . . . . . . . . . . . . . . . . . . . . . . . . 13.245
Root chorb. i n . . . . . . . . . . . . . . . . . . . . . . . 18.916
A i r f o i ls e c t i o n( p a r a l l e l t o rootchord) . . . . . . . . NACA 654005
Sweepbeck leading edge. deg . . . . . . . . . . . . . . . . . 55
Pod t ~ i l s :
Tots1 zrea(onefin 1 . sq i n . . . . . . . . . . . . . . . . . 70.848
Sezispan. i n . . . . . . . . . . . . . . . . . . . . . . . . 9.600
Rcot chord. i n . . . . . . . . . . . . . . . . . . . . . . . 9.840
Taper r a t i o . . . . . . . . . . . . . . . . . . . . . . . . 0.500
A i r f o i l s e c t i o n( p a r a l l e l t o rootchord) . . . . . . . . NACA 6 5 ~ 0 0 5
SweeDback lesding edge. deg . . . . . . . . . . . . . . . . . 52
Angie between t a i l s . d e g . . . . . . . . . . . . . . . . . . 120
DELT4-WING A I R W ~ MODEL (SEE! FIG. 3) Open nacelle
T
- ~ ~~ Pjacelle statiorr Internal Radius A Radius B Dininsion D Diniension 3 radius
" " _ " " _
-6. OCO ""- ""- ""_
""_
""- ""-
-2.440 ""- ""_
""- " " _
""_
""- ""-
-. 621 ""_ ""- ""_
""- ""-
. 000 " " _ ""_ ""-
" " _
0.135 "-"
950 3.460 ""-
1- 730 1.700
2 . 0 0 0 ""-
1.895 ""_
3.875 1.829
5.000 2.265 1.135 "-" ""_
4.870 1.969
8.000 2.485 ""_
2.485 1.535 2.091 5.555
11.000 2.590 ""-
1.915 2.590 2.113 5.980 i3.300
2.600 2.055 6.130 "-"
2.600 2.113
i
16.000
2.600 2.125 ""-
6.230 2.690 2.113 19.000 2.600 2.203 6.200 2.600 2.600 2.113 22.000 2.600 2.000 6.100 2.600 2.600 2.113 25. ooo 2.580 1. E49 5.913 2.580 2.580 2.066 28. ooo 2.520 1.655 5.6Go 2.520 2.520 1 989
31. coo 2.430 1 . &lo
2.430 5.273 2.430 1.911 Y 3;. 000 2.265 1.130 4.815 2.265 2.265 1.833 2 . 0 5 5 .8hO 37.000 4.273 1.965 1.965 1 756 40. OCO 1.780 530 3.628 1.305 1.678 1- 305
""_
42.000 1.562 3.125 .620 .620 1.62G
- I
Nacelle spiie Distencc G Distance K Ordinate H Ordinete L
- 775 .22 1.00
2 . 0 0 .ago
- 50
.945 ' .69 3.00 4.00 .960 .82 4.12
89 - 350
5.00 5 . 4 0
- 90
5.00
. goo
- 4 r
* . 4
TABLE I11 - FOUR-ENGINE DELTA-WING A1:RPLANE MODEL DIMZIEIONAL DATA
I. Fuselage :
g
Overall. I.ength, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
69.60 p Maximum wj.dth, pod included, in. . . . . . . . . . . . . . . . . . . . . . . . . .
4 .@I VI
bximum height, pod included, in. . . . . . . . . . . . . . . . . . . . . . . . . .
8.06 H Fuselage base area, sq in. . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7.62 $
11. Aerodynamic surfaces: ( a ) Dimensions f o r main surfaces and pod surfaces:
I Main surfaces I Pod surfaces I
I Dimension, u n i t
I Wing1 I Vertical t a i l I Wi.ng 1 Canard 1 Tail. 1
Span, in. . . . . . . . . . 45.49 10.58 1.4.16 7.1.2
Root chord, in. . . . . . . 43.4.0 10,80 13.31
6.79 """""""" """"C " " " I "
c ' , i n . . . . . . . . . . . 28.94
Area, t o t a l , sq in. . . . 90'7.26 80.00 21~. . l a
95.67 """"""" """""""I- .........
Area, exposed, s q in. . . . 68.52
9- 57 NACA airfoil section:
Root t o 3.767 . . . . . . 0003.46-64.069 0003-64 0004.. 5-64 00011.5-64 0004.. 5-611
3.767 t o t i p . . . . . . . 0004.08-63 0005-64 OOOh-. 5-64 OOOJI-. 5-64 oooh .5-6h
Leading-edge sweep, dsg . . 60 60 i 60 60
CC""""""" """"C
Trailing-edge sweep, de$ . . -10 -10 -1.0
Aspect ratio . . . . . . . . 2.10
2.1.0 (Geometric ) 1. IcO 2.10 1.43
Taper r a t i o . . . . . . . . 0 0 . 4 0 0 0.40
""""_
""-""I""" 0 0 Incidence,deg . . . . . . .
I"""" """"""""
Dihedral., deg . . . . . . . 0 0
0 0 0
Twist, deg . . . . . . . . . 0 0
lBor plane or camberedwing.
N N
TABLE II1.- FOUR-ENGINE IN7,W-WING Al:WLANI!: MODEL DIMENSIONAL DATA - Concl-uded
11. Aerodynamic surfaces - Concluded: Plane Carnkered ( b ) Elevons : Area, sq in. . . . . . . . . . . . . . . . . . . . . . . . . .
79.03. 67.60 Span, in. . . . . . . . . . . . . . . . . . . . . . . . . . .
l g . I r l 1)1. 80 noot chord, in. . . . . . . . . . . . . . . . . . . . . . . .
T.-(g 5.79
Root chord location,percent b/2 . . . . . . . . . . . . . . . 14.65 14.63
T i p c h o r d , i n . . . . . . . . . . . . . . . . . . . . . . . . .
0 3 . 2 6 Tip chord location, percent b/2 . . . . . . . . . . . . . . . .
100 79- 70 111. Nacelles (Areas .and diametersgiven are f o r one duct. 1 :
Ovcrall length, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18.667
I n l e t diameter, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.800
:rnlet area, t o t a l , sq in. . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.941~
Spike diameter, nacelle stakion 0, in. . . . . . . . . . . . . . . . . . . . . 0.931~
Spike area, nacelle s La Lion 0, sq in. . . . . . . . . . . . . . . . . . . . . . 0.684
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30 1
S p i k e c o n e a n g l e , d e g
inkt area, net, sq in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.860
Exit internal diameter, in. . . . . . . . . . . . . . . . . . . . . . . . . . . 1.614
Exit internhl area, sq in. . . . . . . . . . . . . . . . . . . . . . . . . . . 2.0116
In3oard Outboard s p l i t s p l i t Siamese
Spanwise location, percenb ~ / 2 . . . . . . . . . . . . )+0.300 64.630 45.72)I
Spanwise location, in. . . . . . . . . . . . . . . . 9.167 14.700 10.400 3
LocaLion of inleL from nose, in. . . . . . . . . . . 23.864 39.576 29.592 F
Angle be Lween chord plane and center l i n e of
z
nacelle . . . . . . . . . . . . . . . . . . . . . . 0 -3 0
Maximum nacelle cross-sectional area, sq in. . . . . 6.08 6.08 13.60 \5: U I H Iu
d
* II I , " (b) Coorlinal e: (a) C3Qrdlnates 01 m e w chard line. N f lrlrfoil ~ c c l l ~ n .
" - . """" "" I " Chordwi ne Typical span st.atlon Typical qarr Nonrlimansional station airfoil ordinates Y .I 12.000 i n , coordinates Y - l2.ooO i n .
I
Percent Nondimcnsional orainate , A, I3 J ordinate In. i n .
2 , In. line1 in.
zh "" "- - " ~ ~ ~ 0.0286 0.3113 0 0 0 0
,236 .oj2 . Ob8
.023l . PI7 . 1 5 h
. oh4
.0196 -257 313 m 352 .
.I166 .120
. 01 G 7 .200 .625 ,035
. O l l k l . i G s ,192 .116
.9% .567 ,1112 1.250 .G51 .134 .0119 .e% ,0090 . n o 2. yx .m * 513 ,105
. Oooo 1.090 - 769 ,221
a036 3.750 .076 5.000 1.230 1.025 .252 ,006j 1 . 4 4 1.530 .298 .00119 .059 7.500 1.620 2. o y .Oll3 10.000 352
. Qo%
,00211 1 5 . w r ) 1.041. 3.076 . Y I O ,029 , OOlh -017 20.000 ,405 1.965
. 000 25.000 2.0211 .1115
. m 7
2.040 . l11R
.m . 0 0 2 50.000
0 0 55.000 2.027 . h16
, l l o D 40. M)O 1. gtw h5.000 1.9211 .595 1.039 .Y17 5 0 . ooo 55.000 1.751 .355 60.000 1.6oh .329 65. o m 1. h58 .599 70. coo 1 . 4 5 ,266 1.117 .229 75.OM) 80.000 .924 . I & 85.000 .1117 e719 90.000 .103 .5m 95.000 .276 * 057 loo. 000 0 L E . radiun .lo3 .030 CB 20.507 CA h s c o l m 16 i n percent I forward Of the polnt of taugency and perccnL T aft o r this point.
-I, NACA RM ~ 5 5 1 2 ~
L Nacelle ordina:es T 5. .ti02 w 10.667 14.667 2.657 5.333 8.003 17.333 ~
u I u 4 u U U U
1.267 1.267 i . 257 1.267 1.255 1.233 1.253 1.252 1.252 1.267 1.267 1.217 1.2LO 1.257 1.260 1.257 1.260 1.207 i . 2co 1.240 1.2m 1.245 1 . r60 1.235 1.167 1.222 1.237 1.217 1.123 1.233 1.085 1.170 1.202 1.233 1.150 1.017 1.013 1 . 1 3 : 1.162 1.123 .948 1.177 1 . ogg 1.152 .858
- 937 1-1-53 1- 077
1.043 .53s 1.113 1.13 1 . ola . 7 ' . 3 .TOE 1.072 1 . 0 7 0
.975! .9b7 - 392
1 . 0 1 7 1.312 .a48
.530 - 093 .337
""- " " _ ""_ ""- ""-
.212 . ?eo
.952 -716 .945
""_ ""_ ""_
" " e .850
- 633 * 635
.543 .368 0 .713 .693
-"" ""-
.478 .k78
""-
0.920 0
- 233
. b6?
.563 1.067
- 532
. L49
2.000 .375 I. 250 2.503 .375 2.750 1.333 .375 2.667 3 . jC0 h. 300
h. 000 - 313
16. 000 4.300 m235 -100 j . 000 17-333 c r. I U 1 "" .
Section A - A
Stotlon I O 667 I Stotion 12 000
Stohon 2 667 Stotlon 4 0 0 0 Stotlon 5 333 Stotlon 6 667 Stotlon 8000 Stohon 9 333 Dlmenslon Dlrnenslon Dlmenslon Dlmenslon Dunension
- " Dlrnenslon- " "_
0 C B C - - B C B C E C B C
" " " -
.219 1.241 .e21 I 2 5 4 .221 1.255 .222 1.257 223 1.261 223 1.264 .223 I 1.264 I .223 I I 2 6 4 " , " 436 457 1.199 .441 I 213 .445 1.223 .448 1.231 .452 1.242 .454 1.248 I .457 I 1 . 2 5 2 I 1.255
-
,715
I 1.239 I .715 I 1.239
~~ 1.009 I 1.203 p.009 I 1.203 _.
1.017 1 3 5 3 I IO80 I ,913 1 1.149 1 .964 1 I 196 I 1.003 I I 2 4 3 I 1.043 I 1.289 I 1.081 1.335 I I I 2 0 I 1.335 I 1.120
1.339 i . 7 7 : T l .OIO i 1.469 i .84Q i 1.535 j .92E
"
- 1 - 1 .726 j 1.256 1.364 I .497 1 1.453 I 3 2 9 I 5 2 6 555 I 1.599 I .502 1 1.674 .609
~ ~ ~ 1.317
I .232 I 1.423 I .251 I 1.517 I .267 I 1595 I .281 I .295 I 1.751 I .309
I 1.673
1.337
I 0 1 1.445 1 0 I I 5 4 1 I O 1 1619 I O I 1.699 I 0 I 1.778 I 0
Sectlon 8 - 8 Stotlon 13 333 Stohon 14667 S totlon 17.333 Dlmenslon Dlmenslon I -
1.166 ] 673 I I I 6 6 I 673
0 .907 1.669 .907 1.641 428 I142 1.161 1 .974 I 1.165 I .978 -067 1.113 I 1.327 1 I Note A l l letter dlmenslons deflned on figure 7 .
.933 .040 1 .SO3 1
1.000 1.045 1.051 I 0 6 7 1.464 1.064 D i b tunre Rsdius ihccllc " " _ . .
.'.Intion c D E 1 3 A "" - " I""" ""-" 0 0 -5.333
" " " _ " " " _
.191 -2. Mxl
. ?43
.>a7 .364 0.083 0.083 .423 .21a .2/0 .335 .6go .6yI .43a .5G -478 .GOT . !p .920 .627 1. l a l . l b 7 -798 Naca l k Nacelle 1 . a72 .773 1.252 conlRur contour .733
""I" ""-
1.?6/ """"
1 .21q ""-
""- """-
.957
""-" "_""
""" .475
""- """- " " " _
.110-/
"_" " " " _ ""I"
.210 " - " " Dletence up
r -r
fmn nacelle ccnler line 16.667 2 h . 000 17-333 i ? 2 . om 2 ) . .333 - - " - "-
_""" ""- ""_ """ ""_ ""_ """ " " _
0.161 0.065
""" " " - 1 " " _ """ 0 . 00'1
""" _"" ""..
.083 """ """ -"" """ """ .117 .500
""" ""- ""_ """ """
.1L5 .
""_ """ """
""" ""- A 3 3 .J77 ""_ """ """
""" ""_
1 . W Y .m
""" ""_ " " _ """ -"" """
1.167
""" " " _ """
""" ""- 0. a27
1.335
" " _ """
""" 1. s o j o o 0 . ~ ~ 1 3 .063
""" """ " " _ """
1.667 .o52 ,102 1 " " """ .142
1 . 0 5 3 - 093
-1 " - " " . - " -.- " " -
D I c r r Landing-gear VI
1 " . "_ "
UI H Iu
I
I Fuselage \ I/ P a r t 1 Pod Figure 1.- Sketch of two-engine dclLa-wing airplane model and wind-tunnel sting assembly. See table I for additional. dimensions.
"
L-7 83 6 1
Figure 2.- T'hrcelquarter front view of t h e two-engine delta-wing airplane model mounted i n the Langley 16-fool; transonic tunnel.
m . t 23.212 7.850
Dist.'G - Dist.'G -
Ord. H Ord. H " " 1.70 1.70 42.000 0.000 5.000 I1.000 16.000 22000 28.000 34.000 4 Nacelle S to tic
"iF; 1.600 1.2'70 thrust
.04 7 View A - A Nacelle rake No scale No scale Typical section No scale See table XI for nacelle and nacelle spike dimensions.
Figure 3.- Nacelle and nacelle spike configuration f o r the two-engine delta-wing airplane model.
L Pod I Figure 4. - Four-engine &e;'ta-wing airplane model with split mce].les.
See table I11 for additionalinformation.(All dimensions are i n inches. ) s I I , .
Figure 5.- Details of the cambered-wing construction f o r the four-engine del.ta-wing airplane model. See table I V for dimensions not given on the figure. ( A l l dimensions are i n inches.
Figure 6. - Geometry or inboard s p l i t nacelle and nacelle strut. See
table V f o r dhensionsnot given on thefigure. (A11 dimensions are i n inches. ) , I I -. .c .- 4
_"" --"
- _ "
~ \ . / -&"" I
- - .- - - y - < : " "-Y " - ---
""
\
I
I - Sallcm &A SCCIID~ 8-D Srclmn C-C Nncellr slnt~on Nawllc 51aImn Typlcol for n o d l o 5ecllon 0-0 1101m5 0.12MX) I3 353 14667 Nnrellf slallon 1 7 333
Figure 7. - Details of outboard eplil; nacelle. See table V I for dimen-
sions not given on the figure. ( A l l dimensions are i n inches.
w w t
r' h
Figure 8. - Sketch of Siamese nacelle and strut. See table VI1 Tor dimnen-
sions not given on the figure. (All dimensions are in inches.
L NACA RM L55127b
L-81582
(b ) Cambered wing; Siamese nacelles.
Figure 9.- Tlvee-quarter front views of thefour-enginedelta-wing air- ?laze model mounted i n t h e -Langley 16-1"oot transonic tunnel test section.
36 - NACA RM L55127b
L - 81732
(b 1 Cmkered wing; Siaxese nacelles. L-81581
Tigu-re 10.- Front views of the four-enginedelta-wingairplanenodel mounted i n +,he Langley 16-foot transonic tunnel test section.
W C A R M L55127b
(a) Pod wing mounted on pod. L-81735.1
(S) Pod wingrrolmted flush and f a i r e d i n t o u n d e r s u r f m e of mzinwing (necelle rexoved 1 .
NACA ~551277~ I 3 I I IO
. 7 .8 .9 I .o I,I
Mach Number, M Figure 12.- Vzria5ion of Reynolds rmker w i t h k c h number f o r the two- engire End four-engine delta-wing airplane models i n t3e k n g l e y ~ ~ 6 - f o o t transonic tumel.
.oo I Q c - . I .8 .9 I .o 1 . 1 Mach number, M Figure 1 3 . - Effect of k c h number on Fod base-press-me coefficient and iaternal-force ccefficient f o r two-er_ginedelta-wing airplane model.
* !- . "
, , " " , . " I
-1
I C Fuselage c"- - ---- 2Smmese nacelles 2 lnbomd 5 ~ l 1 t nacelles
ip
Angle of allock,a, deq Figure 14.- Variation of base-force coefficient, internal-force coeffi- cient, and point mass-flow r a t i o with angle of a t t a w A x a range of Mach numbers. Four-enginedelta-wingairplane model with plane wing.
6 = oo.
c " _ - b 1 I04 IO0 a5 (a> 6 = oo.
Pigure 1 3 . - Aerodynamic characterj.s.tics of the four-engine delta-wing air-
plane models (plane and cambered wing) with split nacelles. Flagged symbols indicate cambered-wing model.
Lift coeffrcient , C , (b) 6 = -2'.
Figure 1 . 5 . - Continued.
I I M I O 6 I04 U ' 97 $ U
95 g
c U 93 a,_ OI c t a5 .70 * Lift coefficlent, C , ( c ) 6 = -hO.
Figure 1 5 . - Concluded.
Lilt coefflclenl, CL (a ) Cambered wing and plane wing. 6 = 0 ' .
P'igure 16.- Aerodynamic characteristics of the four-engine delta-wing air- plane models with Siamese nacelles.
, .
I ?
VI UI H Iv M
d
I 06 I 0 0 43 %
.70 -
(b) Cambered wing, 6 = - 2 ' .
Figure 1 6 . - Continued.
Llft coefflcient, C I .
( c ) Cambered wing, 6 = - 4 ’ .
P i m e 16. - Concluded.
I a 1
T
F
(VI IO4 IO0 ' 3 7 a5 .90 . 7 0 Lift cooffrcrenl, CL
Figure 17. - Aerodynam:Lc characterlstics 0% the four-engine delta-wing air-
plane models (cambered wing and plane wing) withoutnacelles. 6 = 0 ' .
Flagged symbols indicate cambered-wing model.
Figure 1 8 . - Aerodynamic characteristics of the Your-engine delta-wing
airplane model with plane wing and outboard nacelles off'. 6 = 0 ' .
L a I I I - 2 - I 0 1 2 - 3 4 L i f t coefficient, GI- Figure 19.- Aerodynamic characteristics o:t t h e four-enginedelta-wing air- plane model. with plane-wing s p l i t nacelles Tor several pod modifications.
6 = oo.
u l 70 -L 0 Lift coefflcient, C, (b) Pod wing faired into undersurface of main wing.
Figure 1 9 . - Corltinued.
I 1 I
c
Figure 1 9 . - ConLinued.
h i IO5 I 0 4 70 --* LIft coefliclent, CL (a) Pad wing, ventral fin, and canard off.
Figure 1 9 . - Concluded.
- 1 .
I r . ” .05 - Cornbered-Siamese .04
Cambered-split -+
“ I I 1 - .03 ” - ”
-
”- .02 .7 .8 9 IO 1.1 Mach number, M (a ) CD against M.
Figure 20.- Variation of drag coefficient with Mach number and cross- P sectional area diagrams f o r the four-engine and two-engine delta-wing L N airplane models and model 1 (ref. 1 ) . 6 = Oo.
.
Ln e Frociion of length , x / z (b 1 Area d5agrams.
Figure 2C. - Zoncluded.
. . "" I c I E 0 4 .O?
Pod wing and -
I?
UI ventral fin o f f UI H
Pod wing , ventral fin - I
d"
.Oi E 0 ,01 c C ! !
0.
.c Y- a l
I I
Much number, M (a) CD against M .
Figure 21.- Variation of drag coefficient with Mach number and cross- sectional area diagrams o f ' the four-engine delta-wing airplane model with plane-wing split nacelles for various pod modifications. 6 = Oo.
C z d 1 b .
-. - I 1 I c .O .a .9 I .o .7 .a .9 1 . 0 1 . 1 Mach number, M (a 1 CD agcinst M.
Figure 22.- Varia-Lion of drag coefficient with Mach number and cross- sectional area diagrams of thefour-engine del.ta-whg airplane model with plane wing f o r severalnacelleconfigurations. 6 = 0 ' .
Fraction of length, x / l (b) Area diagrams, Figure 22. - Concluded.
-008 .004 Q1 Q E (b ) Split nacelles.
Figire 23.- Effect of Yech nmker on the redxction in drag coefficient due t o cz.xber a t lift coei'ficients of 0 . 1 3 , 0.20, and 0.25 for the four-enginedelta-wingairplanexodels. 6 = Oo.
- XACA R M ~ 5 5 1 2 ~
.3 3 .02 .o I I 04 + K % V .- 't .02 aJ [JI
e
[3 .01 n v .03 .02 .01 F'igure 24.- Eflect of canber on drag due t o l i f t a t a Mzch number of' 0.90 f o r the tow-engicedel+,a-vingairplaliemdels. 6 = Oo. (Flagged syn- Gols are f o r cankered wing. ) . 8 - 9 I . O
Mach number , M
Figure 25.- VEriation of maximum l i f t - d r a g retio ar,d l i f t c o e f f i c i e n t f o r Tzxirnm l i f t - d r a g r e t i o with lkch number f o r the four-engir-e delte--wing airpleneEodels. 6 = 0 ' .
.
wing- ndcet res' I l l
- Cambered-split
""-
Cambered-Siamese
"-
- Plane - split
. 0 6 .OZ Plane .04 Cambered .04 .8 . 9 I.o I. I Mach number, M Figme 26.- Vzriztion of the lift-curve slopes w i t h Yach nuiber for the ?ox-enginedel+,a-wing airplane models. 6 = Oo.
" I L
Plane - split Plane - Slamese
I ..
. 7 .8 . 9 1.0 I. I .7 .8 . 9 I .o 1 . 1 Cambered - 9 o m e s e
Cambered - split
Mach number, M
Figure 27. - Ef.I?ect of Machnumberon the zero-Lift pitching-moment coef-
i'ic3.en-k f o r the four-enginedelta-wing airplane models. 6 = 0'.
c . 3 I I I I I I
Wing - nacelles
Cambered - s p l i t
"""" Cambered -Siamese
- "- -
Cam bered - none
"-
Plane - split
"" Plane -Siamese
" -
Plane - none
Note: Moment center located at 0.35 E c - * ' . 7 . 8 .9 1.0 1 . 1 Mach number, M F i g z e 23.- Effect of Bhch number on the s1oI;e of pitching-moment coef- riciect against lift coefficient for the four-engine delte-wing air- models. 6 = 0 ' .
plane
T
Plane -Siamese PIane-split
Cam bered - split Cambered -S~arnese
Mach number, M f Figure 23.- Effect of Nach zumnber and elevon setting on the pitching- coxent coefficient a t e l i f t coefi'icient o f 0.25 for several configu- rations 03 the four-er?gine delta-wing airplane rmdels.
NACA RM ~ 5 5 1 2 ~ r
Wing - nacelles
Cambered - split
" " " _
Cambered - Siamese
"-
Plane - split
Mach number, M Figure 30.- Variation of elevon effectiveness parameters with Mach num- ber for the four-engine delta-wing airplane models.
* *
Wing nacelles
Cambered - split
----------- Cambered-SIamese
Mach number, M Mach number, M Figure 31.- Variation of elevon angle, angle of attack, drag coefficient, 'and lift-dragratio at -trim (maintaining a 3-percent static margin and a lift coefficient of 0.25) with Mach number for the four-engine delta- wing airplane models.