Document
ff 853 July 85 ’ W Introduction mmercial f l i g h t a t supersonic speeds has served fcr t h e
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l a s t deczde 8s a stimulus t o aeronautical research i n t h e United S-cates zr.Z A- c--,. 1 . 7 Within t h e United S t a t e s the intensive development p r o g r a x i n * r . p 2 - d u ~ ~ 7n-4- of su2ersonic military a i r c r a f t provided a broad backgrom.- af z?chnical kcwledge uoon which t o base i n i t i a l concepts, studies, ar,d reszarch 2ro;rms. ‘ 3 2 s p o t e n t i a l i t y of comercial supersonic f l i g h t w a s s p e c i f i c a l l y assessed i n December 1939, i n a technical summary of supersonic t r m s p o r t (ref. 1). Research problem >repared by NASA and presen-ced t o t h e FAA, efZort within NASA and elsewhere increased s i g n i f i c a n t l y with t h e inauguratior, of a national program i n 1961 and continued t o expand as The prograx Z a i n x / ’ . momn-cum.
I;ASA r e s z a c h has exompassed most of t h e basic aisciplines. IL t h e ; - e a of ccriI‘igaation aerodynamics, for example, over 30 basic configuration c ~ x e g t s have been explored. The f o u r most promising concepts emerging i n 1e;e lS62 Two of these c c r 2 i x d a t l m s f r o m t h i s research are i l l u s t r a t e d i n figure 1.
a r e fixed-wing arrangements and the other two incorporate variab le-sweep wings. A recen-c and more completz discussion of t h e four SCAT (Supezsonic Comaercial---ir Transport) configurations i s presented i n referencz 2. I n o r s e r ; a p-ovide a focus f o r further TUSA research, contracts T : C X l e t t o t - e 33eing and Lockheed Companies ir, 3’ebruary 1963, t o study these four from t h e viewpoint of t h e aircraft co?ii,--ation concepts i n depth The studies were completed i n manufzcturer and t h e a i r l i n e user.
Septezber 1963, and were reported t o t h e industry a l m g with other research results i n a supersonic transport conference held a t tl?e Langley 3esec;rch C ~ ? - t e l r on Septw.ber/g-21, 1963.
The brozd conclilsions of t h e SCAT F e a s i b i l i t y Studies and “,:?e ac;smpanin, research studies were as follows. Derivatives of a t l e a s t two of t k e ? o w c o n f i p r a t i o n s were judged t o be technically feasible i n t h a t they co;lld E G e C U zhe b w i c rxission requirements within t h e prescribed operating r e s t r i c t i o n s .
-*-‘ These studies derr;onstr&ed t h e desire-aility of a T i t a n i u m airframe and t h e Eowever, it TTZS indicated t h a t t h e resulting necessity f o r dvanced engines.
and a i r n l a e s would be l a r g e r and heavier than corresponding subsonic jets, It was obvious t h a t ways would t h e i r econonic f e a s i b i l i t y was questionable.
have t o be found t o obtain f u r t h e r major increases i n f l i g h t efficiency.
I-i ~ 2 s c l e a r t h a t major attention would have t o be paid t o t h e sonlc boom, .
T L - X =- W ~ S shown t o have become a dominant f a c t o r i n a i r c r a f t design m d oper a t ion.
Tne purpose of t h e present pzper i s t o review some of the major reseacl= a c t i v i t i e s of t h e last two years, conducted by IU’ASA, i t s contractors, 2nd i n su2port of t h e supersonic zransport with emphasis i n t h e areas o t h s s of iq3raved f l i g h t efficiency, s t a b i l i t y m6 control, structures and naterials, snd operati26 poblems.
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Kast td:iy s q e r s o n i c research ained a t drag reduction dealt vizr- wave Gre,.
These ini'iial e f f o r t s led t o t h e ''area rule" concept, and t k c u g h ex'censfve calculative and experimental research, t o t h e present l i n e a r theory nezks&s f o r t r e a t i n g wave drag and compocent interference.
These a n a l p i c r?.eti-ods generally show good zgreement with experiment for rzesonzkly slender confizuratio3s.
Program hwe been developed f o r high-speed computers tsihich have zwned i i n e a r theory methods i n t o a practicable and powerfa1 deEi-2 tool.
I n the area of supersonic drag due t o l i f t , experiment shovea -,k.,zt tlx theC4re-cical Lains were attainable only f o r conditions where l i n e a r theory was G2:jLicaLle - t h a t is, where t h e configurations were reasonably slender arid i o c a l S ~ C Q ~ S not extrece and where cruise l i f t coefficien-is vere r,oderately ~ G W .
The Sorementioned progress i n t h e area of wave drag and interference le& t o the lcv-drag, slender corfigurations for t h e supersonic transport srhLck rret t h e linear theory requirenents.
L i n e a theory methoas were dL-;e:G-;?d Tor hmdling a wide variety or" arbitrary planforns, ail2 zxserimen-, sLnyT . , d as muc:-- as 85 percent of t h e thzeoretical inproverrent i n d r q - & e - t o - - .
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du? t o t w i s t and caa'oer could be o5tained for arrow t E e planform.
k ;et 0: conputer programs was developed t o mechanize tkiise complex co:.x-,uz,:ions. The resulting 2rogr&xs, x i t h i a the r e s t r a i c t s oI" lir,ear o f cam-Der surfaces of a r b i t r a r y wing planforms theory, 2ermit calculations with ssecified pressure distribGtions, Inversely, they permit deterrninatioc GI" s r e s s u r e Ciszributions on wir.g pianforzs with a r b i t r a r y surface war2ing.
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'I . . 'I - 2 - X g - u ? 2 i l l u s t r a t e s a t y i i c a l build-ur, oI' a configurazion drag 2 0 1 ~ Items shown a r e wave 2s ~ c c c r ~ ~ 1 T s k a i by use c? t h e XASA computer programs.
-.. Gd, " ,-- f'1-ictiz-. drag, and drag-due-to-lilt values f o r : (1) t h e selectzed T.T^-- ..Ld2~c-vmg .
emfiguration, (2) a f lat-wing version of t h a t same configuatio-, a ~ d ( 3 ) t h e t h e o r e t l c a l lower bound for t h e chosen wing planform. Part ( 3 ) , the lover h w d , provides a quick azsrrer as t o ~ ~ ~ h e t h e r o r not the aerodynamic Any prcperly perfomence L j e c t i v e s can be met v i t ' n t h e plarxform selected.
ilrarped wing must f a l l between t h e flat wing and the lower bound, but 1,511 n o t be zecessarily tzqgent t o the lover bound because of p r a c t i c a l design r e s t $ p n t s .
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- 3 - Tkese calculative rnethods which provide t h e drag polars of f i s r e 2, .
-.-hi ch ilivolve t 3 e a e t e m i n a t i o n of c&oer-pla??e pressure d i s t r i b u t i o r s such a3 l i f - c - LLT- czn 3e used 'io develop other aercd7pmic characteristics, c r v e slope, pitching moment a t zero l i f t , s t a t i c longitudinal s t a b i l i t y ani wine loads. T h L o information con be grovided not only for the cruise condition 'aut ;"or a ranze of lift coefficients axd susersonic Kach nuabers.
A :,:ar?ed-ving, Supersonic-transport -type configuration developed through z;ie use of i;kese new calculative methods and! designed t o cruise a t Mach The wing leading-edge i s swept behind t h e nurrber 2.6 i : , shown i n figure j .
- 7 - L - L - I I L x h sor-e/ ,he engine nacelles have been loczted under the ving and t o t h e 'irne f i n a l rezr,.;z a 2 o s i t i o n Yavorable for l i f t m d Crag interference.
confi->,:.-.ation vTas developed from t h e i n i t i d concept by progressive? i.t?rat:c: by LZZ~:S o f t h e c c q u t e r prograns. Zxtensive vind tunnel t e s t s vere corduct& t o ckeck t h e v z l i d i t y of t h e computer r e s u l t s 2nd t o evaluate t k e coEllguzztio:i.
The experiEenta.1 points p l o t t e d on figure 2 demonstrate the excellent agreement t h a ' : c z n be o'stair,ed 'oettreen e x p e r b e c t znd theory.
It i s stressed tk'; t h e s a q l e configuration shown i s one of a number of -'ix&-wlng end variable-wee? arrangexents being studied by NASA t o correlate ex-)e-.:rr?:T,t and theory and define the z.erodTJr,mic s t a t e of the art. The r;iL":2ose of this work i s t o dernonstrzte t h e c a p a b i l i t i e s of new aerodynanic t o o l s =xi techxiques r a t h e r than define a s p e c i f i c supersonic transport conffipration. Xmy other fectors, such as take-off and landing c h a r a c t e r i s t i c s , zener21 arrz:~ecia-~t, s t r u c t u r a l f e a s i b i l i t y , and ease o? f a b r i c a t i o n must be ccr.iidc:-ed i n C e t a i l by qualified aircraTt designers a d s i r l i n e operetors belori-- .I opzkiun configuration can be selected.
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A i t e q t s 'io i q w m e YliZht e2f'iciency ca;lr,ot omit consideratic-s 32' t r i n m i y t h e cc:li"5p.raJtAon a t some reasom,ble l e v e l of sta'oility c m s l s t e n t r v.- . i -G;2 x 3 s o n i c rzquirenents and t h e t o t a l c h a g e i n s t a b i l i t y from the lev- , - , , L , G -,-, 2 <> -Uo t h e supersonic cruise Mach number. Tn contrast t o t h e f l a t wiR2 c a s t , . .
. L., .- i s p c s s l j l e t o design t h e v:arped wing with a pressure d i s t r i b u t i o n r,rhich ~ E . : : i . z ~ ~ e s tlie ::zcessl+,y f o r CGnti-Gl deflection a t t h e cruise l i f t c o e f f i c i e z t Tne cats, shorn (ccnfign-s,tion hac a positive pitching mment a t zero l i f t ) .
iLl p i e l L , a ~ e 4 ( a ) compae t h e trilmed l i f t - d r a g r a t i o f o r t h e coqara'sle "I"lat" Tiiese data show t h e somevhat IiiZher peak efficiency :-.e I t - ;=.r,edl' configurations.
- .-. - - > - .
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L1--z'.~T--r ,L-pza I . - : - - : , ccr3pazed t o xhe f l a t wing, and t h e relative i n s e n s i t i v i t y "- --c --,.- ~ - - ~ ~ . r - - - n ~ = \,.i-n.- . _ c . LO s t a t i c aargin as p r a c t i c a l s t a j i l i t y levels are reached.
"his f a c t o r i s G: ma2or significance t o t h e SST i n view oi" t h e wide range of loaCiizc.;5 and sre:-irztion of s t a b i l i t y level with Mach number.
3.. uL.c:x -. ~ . , ' s-iuiies have shorn t h a t the locatior, or" the engine nacelles can As mentioned v-,. v L .,Tr: r; ~ ~ Z : : L L - ? ~ ~ lavorzble 21-q aEd l i f t interference effects.
-----. .-.,< --- i : , c u - ~ l ; r tl:e e z g i x nacelles of the s m p l e configuration vere located waves f r o = t h e nacelles 'oezzztk m d v e l 1 aft on t h e .wing. Tae c o q r e s s i c n L?2inze upon tlie receding slopes of the wing, thereby producing both favorable I n t h i s case these e f f e c t s are lift a d thrus-c interference, f i g u r e b(b).
seen ':.- s'L'3iciently po:rerful t o overcme both t h e ware and f r i c t i o n drKs G< - c L -.aceller T c i - l i f z coefficients a t and above t h e cruise l i f t coefficient.
A E i t i o n a l recsarch oa these interference e f f e c t s i s presently underwzy.
The i:ni;il-c~.;:.t~t i n fliGh?t efficiency (JIB T , ) achievsd i n the period
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SFC betveen - G , : ? SCA2 f e g s i b i l i t y studies a d t h e present data i s shown i n : i s r e 3 .
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-L--2~le /. r;r.i, c x i i g u r a t i o n s just discuszed f a l l s a t t h e top of t h e “AZvanced AA?ro2 7-,? ..~lmics’~ ,- Szrld, an& exhibits the highest level of efI’iciency y& neasured Woke, tm, t h a t t h i s efficiency l e v e l &:proaches - Y 1 . . t:>i _I s Mach number range.
-L-. -, L : L Z G< t h e 2Z t o continuous m d p a r a l l e l experimental and calculative research g r o g r m s a h e d a t irzSroving * su3ersonic performame.
h y cocclusiozs 2s the o v e r a l l vehicle eZ’iciency must be based on en , Complete mission ax,-,-,rsis o l : %!?e c o q l e t e mission frcs take-off t o landing. -J studies are 811 &solute m c e s s i t y t o deterair-e t h e performance p o t e n t i a l G ? a ~ i v e n :;i--9lane, t o define problem areas requiring research, erd t o . . ..
* The e f f e c t s of design cc -.zctly e v a h a t e t h e e f f e c t s of design changes.
chz;.1c:.es c a x o t be deteriiined on t h e b a s i s of individual conponent ckaracter- i s t i c s beczJse of t h e corr:plex interzctions which exist betveen <he sirfrzz:?
eerot:r:xiics2 p - o p l s i o n system, weight, and operat in2 requirerxent s.
- -- . : I -l-G -.a s q e r s o n i c trensport perfomance i s currently leterrnixed by t h e cse Such program are more o r less a? s o l 2 i s t i c a z e d nachi3.e coqmter program.
Using s p e c i f i c airplane and engine character- s t e ~ k z r d ?;::youghcut inzustry.
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-u,,.-cs - .,-- t h e programs a r e designed t o s e l e c t f l i g h t paths which s a t i s f y greselected ace-leration ar?d cruise sonic boo2 l i m i t s , t o s a t i s f y engine c;:?-k-;ional l i L : i t s , to s e l e c t t h e cruise altikude which vi11 r e s u l t i n m a i m m .- -L:--:z, .~ t o l i m i t normal acceleration, etc. The program usec Sy NASA have suff’icien’; f l e x i b i l i t y t o enz’ale the mission t o be studied i n d e t a i l .
' 1 f Conpai*ed t o the su'oso??ic t r a c s p o r t the major new f a c t o r i n tkt: 6ST - .
c . E : : - L ~ process i s the i-ec_uirerr;ent t o hold t h e sonic boo3 overpressures below -,DecL;'" i l e d k v e l s 2 ~ i ~ i ~ transonic x c e l e r z t i o n and supersonic cruise. The bosm levels speciflea &re o f c r i t i c a l LQportance because o f the powerful Figure 6 shows The gross weight for several eff'ects on z.ircraft s i z e .
con2i;urations as 2 function of t h e c?esign m x i m u m overpressure for transonic accelzratioc. Ezch curve represents 8 f & l y of airplanes, each of sLn.ich L., - m e 'seen csirefully matched and optiinized f o r minimm gross weight ZL i t s As the mw;imm allowable overpressurz i s pnr;r-;icuLar overpressure l i m i t .
reCkce5 tke aircraft i s requires t o accelerzte a t progressively h i g k r al-litudes, 7:hich requires greater f u e l consmption, a l a r g e r wirig> azd 11;rger lnes - a i l of which increase the e i r c r z f t gross weight t o c c c m p l i s h t h e - t h e growth I J - o ~ . -2 G.C> case below soze l i c i t i n g value of overpressure, Obviously f x t o _ - prc;ce.;s takes GVZ and grcsc w i g h t increases precipitiously.
+ 7 ? . - U A + L ~- cesirc'-le UJ desi.-r 61 point i s near t h e knee of t h e curve.
r-. ---= . - c j j e c t i v e o f research i n the area of corfigurstion e f f e c t s i s t o This i s accornplished s h i f t the k x e of the curve doimward m d t o t h e l e f t .
n r a c - l i c a l l y : 3 y i;;creasii?g t h e a i r c r & f t f l i g h t efficiency i n a l l p a r t s of t h e E l i 2 s i ' 2 ; : i q r o v i r g t h e sonic 'GJOX f o r m f a c t o r (as w i l l be discussed l-." cbe:-,;. \ C~:=ves A . and B are conI'igurations studies i n t k e S C N f e a s i b i l i t y - .
s-"-;L>z j , . 7 \ ) : - ~ - z a s 7 . - t h s c 2 l e configuretion h ~ s Seen ;"ur",er oFtinisz,5 frori , . L L ? G zt-T-'.r.,-.< "' - u2.,---b of EerocyczzIc efficiency and sonic boom form fector cti1iziL-i .L>- L--, ::ei;ly avz.lla'3le teckiicpes. Taus> sonic boom i s c r i t i c a l l y importznt '3 i n t h i s f i g u r e t o rule out one configuration i n I'zvor of snothel-, ;Lel:r::L.ing 011 < : l e s e t t i n g of t>e zmimm allo\,rzble boom o v e - p - e s s u z .
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- 7 - ~~~~~~~ of ikz *L20rtance, sonic boom resecrch, 50th t h e o r e t i c a l a d c.:prixatal, i x ; I - ==,n 3 vigorously pursued over the 1ss-t re-:: yezrs. 4-z -7 _.
1-"5- ", for li2hte-c a i r c r a f t t k e calculztive as vel1 as tiie expzrine:ltal
X'ajor scL::-55 d e z l t only v i t h ttvolume" effects of t h e configuration.
InC-re - i . > L > - - 3 - :-,- . s i r e alid TreiL;hnt f o r the supersonic transport, hovever, have necccs:ta'ced consideration of " l i f t eTfects" i n t h e case o f t h e SS?.
F;-,, previousiy * - - - - -lcn-iiioned dreg &us t o i i T t p r o g r m have played ~7 inpor'imt II r o l e i n t h e zialysis of such l i f t effects by providing a rapid .malyLlcal - .
This program has been ZZES Tor c.~;ziizinz the required iift Ciistributions.
_ r .-... - - J T , 4 . - - L.v-i-.Ac~ n . , - . ~ A ~ p a t s of -;he Wave d r q p r o u r m t o obtain a new p r o g r m which c, c : - k - ~ k % e s :Le sonic boox shape f a c t o r as a function of lilt coeTTicient.
, 3 2 2 L ~ y o j r z x s are i~,ow being u t i l i z e d throughout industry.
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evslcijne2t of the znalytic c?et2ods ':.T~s accoqanieci by a pti,ellel expc?rLTei;tal research prograa requiring specie1 techniques. I n order t o I .
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aciiieve i n t h e srind tunnel t'ne required "far-field" conditions ( t h s t 1 3 ,
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L ~ : ~ a s u r e i i ~ e c t of the sonic-boon signatu-? being taken many moZel
- r - L ~ x fro^ <?e m d e l i t ' s e l f ) , the :xLels ked t o be quite sxall, ecs The
- . - I I . , A c: T'ne i l l u s t r a t i o n i n t h e to?
e-senzing ap_>zcatusvery sensicive.
por-cfs:? cf f i L p r e 7 provi6es an iclea of the s i z e of the models used.
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-4 cor:-..--;1Gn of ::xilybical, f l i g h t , and vlnd-tunnel r e s u l t s also shmn L- ~ -- 2iEdi-e 7 kCi:,..,-;?s t h a t t h e methods used a - e r e l i a b l e a ~ $ aclequate f o r j e<-< r,n conic boom overpressure levels. It should be noted t h a t the ~ - ~ h c n t a l ?light t e s t r e s u l t s a r e mean v a h e s as discussed i n a iate-c s pa2er. The p e d i c t i o n bmds f o r t h e win5 turznel zid . : c C X i G 3 Gf t h e o r e t i c a l r e s u l t s allor%r f o r t k e variations i n aircraI't gross weight end Early r:.ir:cion studies by iLGA and others (ref. 3 and 4) indiccted c r i t i c 2 1 o 1 = c given fli2:I:C the ensines sperate oyer g r e z t l y varying ccnditions, a wide - c r y - - - d . ~ - L 0 : ' d e s i z : ~ p z l - a e t e r s must be considered, and t h e engine selection nust 7 7 3% rCEL.2 on : n e sssis o f overall xission p e r f o n m c e .
A recent analysis of - Lhis -ky:2e i s ciiscussed i n the a r t i c l e 'oy IJI. Dugan i n t h i s issue.
;<is resiJlts - 2 - ' I .
I ' 7 t I The achievement of a l a r g e r r a t i o of Cn t o C2 has been one goal of
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The e f f e c t of wind-tunnel tests at both subsonic and supersonic speeds.
v e r t i c a l - t a i l position and fuselage forebody cross section on t h e v a r i a t i o n I of C w i t h angle of attack is shown i n figure 9. Replacing the center v e r t i c a l t a i l with two t a i l s having t h e same t o t a l t a i l valume and located well outboard on t h e wing resulted i n improvement in Cn Although a t high angles of attack.
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not shown, C2' was reduced somewhat as a r e s u l t of the reduced height of t h e
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twin t a i l s . Similar e f f e c t s at supersonic speeds have been shown previously .
i n reference 7. Regarding fuselage forebody e f f e c t s , it has been found t h a t s l i g h t deviation from c i r c u l a r cross section can result i n considerable -_. ..
Inasmuch as fuselage improvements i n Cn a t moderate and high angles of attack.
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cross-flow c h a r a c t e r i s t i c s are knownto be s e n s i t i v e t o Reynolds number, research i s currently underwey t o determine if these differences are maintained
t o Reynolds numbers beyond the maximum t e s t value of 6 x 10 6 (based on fuselage
depth).
Inasmuch as t h e general c l a s s of l a t e r a l - d i r e c t i o n a l handling q u a l i t i e s problems discussed herein a r e most apparent at high angles of attack, they are of importance t o the supersonic transport a i r c r a f t i n t h e landing approach phase of the transport mission. hctensive a n a l y t i c a l and simulator studies of landing approach handling q u a l i t i e s lead t o a t e n t a t i v e conclusion t h a t t h e variable-wing-sweep configurations which 'can approach at low angle of attack by v i r t u e of high aspect r a t i o a n d l o w wing sweep angle and u t i l i z i n g e f f e c t i v e h i g h - l i f t systems, exhibit approach handling q u a l i t i e s nearly comparable t o e x i s t i n g subsonic jets. "he fixed wing configurations, characterized by t h e i r high leading-edge weep and low aspect r a t i o and with t h e i r less e f f e c t i v e I I h i g h - l i f t systems, must approach at higher angles of attack and require s t a b i l i t y augmentation with r e l a t i v e l y high authority t o achieve acceptable handling q u a l i t i e s i n t h e landing approach.
Turning now t o longitudinal s t a b i l i t y the rearward s h i f t of t h e aero- cynamic center with Mach rimer coupled with the requirement f o r positive longitudinal s t a b i l i t y at subsonic speeds r e s u l t s i n levels of t r i m drag which could seriously l i m i t t h e performance.
As shown earlier (fig. 4(a)) t h i s penalty can be reduced by t h e use of wing camber and t w i s t .
However, f o r configurations incorporating variable-sweep, variations i n aerodynamic wing sweep angle as well as f r o m t h e character- center r e s u l t from changes i n ~ *..
I i s t i c changes associated with Mach number.
Extensive experimental and a n a l y t i c a l studies by NASA and substantiated by industry studies haye indicated t h a t use of an outboard-pivot location will allow a lower minimum sweep angle while r e t a i n i n g an aerodynamic center location compatible with t h e cruise design point (fig. lo).
A r a t h e r complete discussion of t h e l i f t d i s t r i b u t i o n s associated with the various piyot locations and the explanations of t h e i r aerodynamic center v a r i a t i o n s with sweep angle i s presented i n reference 8.
A t supersonic speeds t h e combination of large sweep angles and high-panel aspect r a t i o s used on yariable-sweep configurations can r e s u l t i n aeroelastic effects s u f f i c i e n t l y large t o provide reductions i n s t a b i l i t y t o t h e degree t h a t t h e supersonic condition could become t h e c r i t i c a l s t a b i l i t y case. It must b e remembered, however, that a reduction i n s t a b i l i t y due t o aeroelastic e f f e c t s does not necessarily imply a reduction i n cruise trim drag.
The cruise teim drag i s dependent upon t h e wing-body center-of-pressure location which i s d i c t a t e d by the wing warp required f o r optimum cruise performance and the f l e x i b l e wing must be b u i l t so t h a t it assumes t h e design shape i n one g f l i g h t a t t h e design speed.
' I i The avoidance of undesirable motions o r excursions i n angle of' attack (pitch-up) requires l i n e a r i t y i n the pitching-moment variation with angle of attack. Past experimental studies have shown t h a t t h e degree of U n e a r i t y i s a W c t i o n of wing-sweep angle, aspect r a t i o , and h o r i z o n t a l - t a i l location, along with other f a c t o r s such as fuselage size, engine location, etc. The fixed arrow wing (e.g., f i g . 3) which is desirable from a performance stand- point, has an undesirable nonlinearity. Research currently underway, however, has indicated t h a t leading-edge devices t a i l o r e d t o a specific configuration can provide a s i g n i f i c a n t reduction i n t h e s e v e r i t y of t h e i n s t a b i l i t y .
For variable-sweep configurations, i n addition t o the c l a s s i c pitch-up -....
associated with t h e high-sweep condition, a second type of pitch-up occurs when t h e outer panel is unswept. A leading-edge yortex, shed from the highly swept forewing, causes t h e forewing t o carry a greater proportion of t h e t o t a l lift a t high angle of attack and promotes earlier stall of t h e outer panel with both e f f e c t s contributing to an unstable tendency a t high angle of attack.
A considerable improvement i n t h e pitching-moment c h a r a c t e r i s t i c s can be &) and obtained with f a i r l y moderate reduction i n inboard forewing sweep as shown i n figure 1 1 , indicating t h a t t h e r e m a y be some desirable.
area, and s t a b i l i t y . A f u r t h e r improvement can be compromise between performance
(h) on t h e highly swept fixed
r e a l i z e d by deflecting a leading-edge f l a p The d a t a j u s t discussed were obtained from a configuration portion of t h e w i n g .
As shown i n figure ll, r a i s i n g the t h a t u t i l i z e d a low horizontal t a i l .
h o r i z o n t a l t a i l as may be dictated by engine location considerations r e s u l t s i n t h e t a i l contributing t o high-lift i n s t a b i l i t y . '%nis may be avoided f o r Additional moderate t a i l heights by, incorporating negative t a i l dihedral.
in pitching-moment. l i n e a r i t y are shown t o be attainable by t h e proper g a i n s c . .
- ..
I application of wing leading-edge devices. Further research t o obtain .
.
more effective solutions i s underway, Operation on the so called "backside" of t h e th-rust required curve means t h a t an increase i n t h r u s t is required t o maintain the f l i g h t path w i t h decreasing speed. Operation a t t h e speed w e l l below t h a t f o r n e u t r a l speed s t a b i l i t y is objectionable. A small amount of speed i n s t a b i l i t y may be tolerable, however, the f i n a l answer t o t h i s question must await the r e s u l t s of f'urther research. For variable-sweep a i r c r a f t speed s t a b i l i t y i s generally assured because c r f t h e higher l i f t coefficients and l i f t - d r a g r a t i o s provided by t h e higher aspect r a t i o s and the a b i l i t y t o e f f e c t i v e l y utiU.ze h i g h - l i f t ?....
* f l a p systems. The severity of t h e speed s t a b i l i t y problem f o r low-aspect- r a t i o fixed w i n g configurations brought about by t h e high induced drag, the inefficiency of h i g h - l i f t systems on this type of wing, and the d i f f i c u l t y i n trimming out f l a p pitching-moment coefficients i s i l l u s t r a t e d i n figure 12.
For the wing loadings shown, landing approach speeds below 170 knots can only be attained through the use of larger wing-flap deflections which would require increased trim capability such as m a y be provided by a large canard o r a rear tail. For t h e tailless delta configuration shown i n figure 12, t h e use of the extremely low+wing-loadings required t o reduce t h e speed corres- i s generally inconsistent with t h e require- ponding t o neutral speed s t a b i l i t y of supersonic performance, and t h e designer may be forced t o accept some ments degree of speed i n s t a b i l i t y .
at D I 1L
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The supersonic transport poses many structures and materials problems associated w i t h t h e long life tine and r e l a t i v e l y high temperatures t o which that may e x i s t i n t h e structure w i l l be subjected. Nonunifom temperatures t h e wing and fuselage sections w i l l produce thermal s t r e s s e s t h a t must be Studies t h a t haye been m a d e on t h e influence of considered i n t h e design.
thermal s t r e s s e s on t h e weight of' the a i r c r a f t structure indicate t h a t weight p e n a l t i e s of a f e w percent of t h e basic s t r u c t u r a l weight m a y be a t t r i b u t e d t o t h i s factor.
Studies of d i f f e r e n t s t r u c t u r a l concepts f o r the supersonic transport i n -%heMach 2 . 5 and 3 range by both industry and NASA indicate that the l i g h t e s t weight s t r u c t u r e s can be achieved through the use of titanium alloys.
The application of titanium favors skin-stringer construction f o r both wing and fbselage structures, whereas, s t a i n l e s s s t e e l would require more complex l o c a l s t i f f e n i n g (e. g., sandwich construction) t o develop high compressiye and t o r s i o n a l strengths.
Among the titanium alloys t h e 8~1-Uo-1v a l l o y is t h e p r i n c i p a l material of i n t e r e s t . This a l l o y is r e l a t i v e l y new and l i t t l e fabrication experience e x i s t s . I n order t o explore some of t h e f a b r i c a t i o n problems associated with s t r u c t u r a l applications of t h i s alloy, NASA has i n i t i a t e d a cmparative study of seyeral d i f f e r e n t methods of fabricating skin-stringer panels representative of wing compression cover skins. Some of t h e preliminary compressive strength r e s u l t s based on an average of three test panels f o r each of t h e three methods of construction are shown i n figure 13. The highest compressive strength was u , ./ I I
. c
- 2 - obtained from t h e resistance spot-welded panels, followed by t h e riveted The maximum difference and the fusion welded (tungsten i n e r t gas, TIG) panels.
i n compressiye strength between the spot welded and fusion welded panels was approximately I 2 percent. These i n i t i a l r e s u l t s were obtained with panels from 8 ~ - N O - 1 v titanium a l l o y i n t h e t r i p l e x annealed condition. Additional panels are being fabricated from the titanium a l l o y i n t h e double-annealed condition using t h e previously noted f a b r i c a t i o n methods i n addition t o electron beam welding, diffusion bonding, arc spot welding, and machining frm t h i c k To date t h e f a b r i c a t i o n problems encountered are t y p i c a l of those plate.
encountered with a r e l a t i v e l y new s t r u c t u r a l material.
Fatigue properties of t h e materials a r e of i n t e r e s t because of the long l i f e requirements of 30,000 t o 50,000 hours. Both NASA and industry studies are underway on the fatigue behavior of several promising materials. The studies at Langley include determination of fatigue strength, determination of r a t e s of crack propagation, and Yesidual strength of sheet materials containing cracks. "his research is discussed i n M r . Raring's a r t i c l e i n t h i s issue.
The requirements f o r long l i f e have triggered many materials research programs t o e s t a b l i s h e f f e c t s of long t h e exposure on material properties.
A t Langley investigations were started approximately three years ago t o study t h e s t a b i l i t y of several t i t a n i u m alloy and s t a i n l e s s steel sheet materials after prolonged exposures a t 5509 f o r times up t o 40,000 hours. Data on t h e 22,000 hours have been obtained t o date. The e f f e c t s explosure e f f e c t s up t o of exposure are determined fram changes i n the mechanical properties a t room temperature as well as -110% and f'rom changes i n t h e t e n s i l e - 3 - .
spot w e l d strength. Metallurgical changes i n the material r e s u l t i n g from t h e exposure are a l s o being studied. The data thus far indicate t h a t the titanium alloys and s t a i n l e s s steels of current i n t e r e s t exhibit no significant change i n mechanical properties. The only clear evidence of deterioration in these materials has been noted i n the t e n s i l e strength of spot welds of t h e candidate titanium alloys and some of t h e s t a i n l e s s steels. The magnitude of t h i s deterioration f o r t h e materials is indicated i n figure 14 i n terms of the r e l a t i v e strengths, t h a t is, t h e r a t i o of t h e strength after exposure t o that before exposure, f o r exposure times up t o 22,000 hours. The shaded areas indfzated the spread obtained $or three titanium alloys; t h e T I - ~ A M M Q - ~ V w a s l e a s t affected by exposure, next was Ti-6Al-4V and then Ti-hAl-3Mo-lV. The spot weld strength f o r t h e s t a i n l e s s s t e e l s ranges above and below t h e data N o general trend f o r t h e steels i s aV8ilable shown f o r t h e titanium alloys.
from t h e tests t o date.
The most important r e s u l t t o date f r o m t h e long time exposure studies has been t h e slow, steady losg of tensile strength of t h e titanium alloy spot welds.
W s result was not clearly established u n t i l approximately 10,000 The r e s u l t s beyond 10,000 hours substantiate hours of exposure had elapsed.
t h e trend. .
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- 4 - For some time it has been recognized t h a t . t i t a n i u m alloys are susceptible t o s a l t stress corrosion a t elevated temperatures and are r e l a t i v e l y immune t o attack at room temperature. The converse s i t u a t i o n applies generally t o o f the national e f f o r t underway t o explore s t a i n l e s s steels. The magnitude a l l aspects of t h e stress-corrosion problem i s discussed i n M r . Raring's a r t i c l e . Langley is studying many aspects of the problem w i t h p a r t i c u l a r emphasis on t h e Ti-8A1-1Mo-lV alloy. It i s recognized t h a t t h e t r u e importance of t h e salt stress corrosion problem w i l l not be established from laboratory tests alone. The actual environment of t h e a i r c r a f t structure i n terms of and retained on t h e structure w i l l have t o be established.
salt%ncountered Experience t o date with a i r c r a f t t h a t contain stressed titanium-alloy parts i n engine areas has shown no d f e i n i t e evidence of salt s t r e s s corrosion.
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The operational boundaries of Mach number and a l t i t u d e within which t h e SST must operate are i l l u s t r a t e d i n figure 15. The limits of this corridor are determined by considerations of maximum lift, buffeting, engine blowout, atmospheric turbulence, maximum temperature, f l u t t e r , boom overpressure, and airframe and engine strength. Such r e s t r i c t i o n s , however, a r e common t o a l l supersonic a i r c r a f t . In addition, because of economics and mission requirements, the SST must adhere t o a r a t h e r r e s t r i c t e d f l i g h t plan. These operational limits and requirements result i n a number of problem areas.
I n o N e r t o define and examine these problem areas, two studies were The first of these studies undertaken by the NASA i n cooperation with t h e FAA.
conducted by NASA Flight Research Center involved a number of f l i g h t s of a supersonic m i l i t a r y a i r c r a f t (North American A5A) i n t o and out of the I;os Angeles terminal area under direction of Air T r a f f i c Control t o examine t h e compatibility o f W e d s q e r s o n i c and subsonic t r a f f i c . A second and more extensive study (ref. 9 ) is currently underway and involves an SST simulator a t the Langley Research Center (LRC) and the a i r t r a f f i c control (ATC) simulator a t t h e National Aviation F a c i l i t i e s Experimental Center (NAFEC) of t h e FAA i n A t l a n t i c City, New Jersey (see figure 16). The SST simulator a t LRC A i s e s s e n t i a l l y a Douglas DC-8 cockpit with current j e t transport equipment i < and with instruments modified where needed t o conform t o scale or range requirements f o r SST performance. The SST sixmilator i s being flown i n the I Simulated radio communications (voice) s t u d i e s by experienced a i r l i n e crews.
and p o s i t i o n data a r e transmitted t o NAFEC by land lines. A t NAFEC, t h e ATC a simulated air t r a f f i c control center operated by simulation consists a E ‘ about 30 experienced air t r a f f t c controllers and simulated a i r t r a f f i c sample.
- 2 - The progrem i s designed t o study a r r i v a l and departure operations of t h e SST t o and from t h e Kennedy International Airport f o r both oversea and domestic operat ions.
Because of the probable handling q u a l i t i e s problem i n t h e approach configuration as introduced i n the section on S t a b i l i t y and Control (see NASA'has undertaken t o study t h e handling q u a l i t i e s of ref. 51, F e a s i b i l i t y studies performed by The several t y p i c a l SST configurations.
Boeing Company, under contract t o NASA, have indicated use of t h e Boeing 367-80 test airplane, which i s equipped w i t h boundary-layer control and thrust Flight tests are mod<htion, would be s u i t a b l e f o r such a simulation program.
I scheduled t o start i n 1965 at t h e Langley Research Center.
Engine noise of t h e current large turbojet transports i s considered t o be of an objectionable level not only i n climbout but a l s o i n the approach and landing when t h e a i r c r a f t are passing over populated areas and following Recent studies of t h e SST have shown t h e normal 2.3O-3.O0 g l i d e slope.
t h a t t h e engine noise of the SST w i l l be of a comparable level. An operational technique f o r a l l e v i a t i n g the noise i n the approach would be t o make t h e approaches at steeper than norrnal glide slope i n order t o increase t h e distance A f l i g h t program was undertaken between the noise source and t h e observer.
t o determine the various characteristics of a i r c r a f t t h a t may l i m i t t h e Tests have thus far been completed on C-47, steepness of the approach.
'11-33, and W-102 aircraf't under simulated instrument conditions ( p i l o t under a hood). Tests using a DC-8F are t o be m a d e i n t h e near future.
I f - 3 -
The .maximum operational glide slope was found t o be about 6 ' f o r both
the C-47 and T-33. For the TF-102, t h e maximum g l i d e slope was a t l e a s t 70, The limiting f a c t o r s were found t o be t h e i n a b i l i t y t h e limit cf t h e t e s t s .
t o increase drag without approaching the propeller windmilling condition on the C-47 and the i n a b i l i t y t o reduce thrust on t h e T-33 without encountering engine flame-out o r appreciably increasing t h e engine response time f o r w a w - off. While the p i l o t s could f l y t h e steeper approaches with only a l i t t l e l e s s precision than t h e normal 3' slope, the p i l o t workload w a s increased.
The introduction of the supersonic transport may be expected t o i n t e n s i f y some of'the noise problems associated with current j e t a i r c r a f t i n addition t o -_. ..
sonic boom. Noise i n t h e introduction of new problems associated w i t h t h e community due t o t h e supersonic transport i s a function of t h e type of power plant used and the manner i n which it is operated, as well as the configuration of t h e a i r c r a f t i n which it i s installed. The manner i n which t h e a i r p o r t noise s i t u a t i o n i s affected by operational procedures is shown i n figure l 7 ( a ) f o r a current intercontinental fan-powered, subsonic jet and f o r some proposed I supersonic transports. 110 PNdb noise-level contours are indicated f o r both Because a i r c r a f t , with t h e origin representing t h e start of take-off r o l l .
of t h e g r e a t e r thrust requirements f o r t h e EST, t h e noise l e v e l s t o the side !
The of the runway are generally higher than those f o r t h e subsonic airplane.
a take-off distance i s generally shorter, however, and the a l t i t u d e over given location i n the community w i l l be generally higher f o r t h e superonic t r a n s p o r t and as a r e s u l t community noise l e v e l s may be comparable t o o r Power cutbacks during less than those of t h e current long-range a i r c r a f t .
used f o r current a i r c r a f t because of noise considerations, i n i t i a l clirnbout a r e and should be an acceptable procedure f o r t h e supersonic transport operations.
I I .- . .
- 4 - The noise during landing approach involves the geometry of the engine For current i n s t a l l a t i o n and the a i r c r a f t c h a r a c t e r i s t i c s i n landing approach.
subsonic airplanes it i s generally agreed t h a t the compressor-fan noise during Landing approach is more objection&le than the exhaust noise. Figure l7(b) shows a comparison of estimated noise l e v e l s during landing approach on a 3 ' g l i d e slope f o r some proposed supersonic transport designs and a fan-powered n e extent of the shading represents t h e variations subsonic transport.
It i s apparent t h a t expected f o r i n l e t suppression of the compressor noise.
will be required t o bring t h e landing noise some compressor noise suppression Current research t o lev&s below those of the current subsonic transport.
of the noise a t t h e source minimize t h e compressor noise consists of reduction of rotor-stator interaction and variations i n i n l e t geometry, involving studies of using a steeper including choking t h e i n l e t . Finally, the p o s s i b i l i t y approach as a means crf increasing the distance between t h e source and t h e observer has been discussed previously.
The noise-induced structural-response problems of the supersonic transport are important from the standpoint of maintaining acceptable cabin noise l e v e l s and minimizing sonic fa;tigue. The boundary-layer noise loading w i l l e x i s t i n o f the f l i g h t . The noise f r o m the engines, f o r nearly t h e e n t i r e duration cqntrast t o boundary-layer noise, i s believed t o be significant f o r only a Sonic f a t i g u e w i l l only be a problem on t h e short period during each mission.
s t r u c t u r e i n the v i c i n i t y and t o the rear of t h e engines. The estimated engine noise spectra peak at lower frequencies and reach higher sound pressure levels than the flow noise spectra (ref. 10 ). Although the acoustic loads a r e more severe than those f o r current a i r c r a f t , t h e design of s t r u c t u r e s t o w i t h - - ' * I I . .
- 5 - .
stand these loads is not a new problem since similar environments have been encountered i n current operational vehicles.
It i s believed t h a t engine noise
I
s t r u c t u r a l response experience t o data i s d i r e c t l y applicable; howewr, t h e I boundary-lqyer noise problem has not been s a t i s f a c t o r i l y defined, p a r t i c u l a r l y f o r long-term exposures a t elevated temperatures.
Both t h e flow noise inputs and associated s t r u c t u r a l responses are currently being studied i n NASA I research programs.
SONIC BOOM The sonic boom discussed previously i n connection with a i r c r a f t design 1 .
requirements, constitutes an operating problem of such importance t h a t it merits special attention. Figure 18 i l l u s t r a t e s some of the basic concepts involved. If t h e shock waves fram an aircraft In supereonic f l i g h t could be , I made visible, they wouU look about like those shown i n the figure. These waves are moving at the speed of the a i r c r a f t and axe &served along shock t h e ground t r a c k and several miles t o each side of the track as transient This pressure disturbances, i l l u s t r a t e d by the I?-wave shape i n the figure.
I
pressure signature has associated with it a 4 which i s a measure of the
in*nsity, and a h which is a measure of the wave length, both of which depend
a Of course, t h e sonic- upon t h e a i r p l w e geometry and i t s operating conditions.
since atmospheric boom signature does not a l w a y s have t h i s N-wave shape, e f f e c t s can cause the peaks t o be accentuated i n some cases and rounded off i n others.
I n i t i a l experiments on the sonic boom, measured under carefully controlled conditions, were begun by N A S A i n 1958 with a f l i g h t test program wherein t h e first sonic-boom pressure measurements a t ground l e v e l were obtained from a i r c r a f t i n sustained supersonic flight. The NASA has performed numerous t h e o r e t i c a l and windtunnel-studies and has worked closely with t h e U. S. Air Force and t h e Federal Aviation Agency i n carrying out flight test programs (ref. U). This research e f f o r t had the dual objective of determining t h e magnitude of overpressure produced on the ground and t o attempt t o e s t a b l i s h t h e t o l e r a b l e l e v e l of sonic boom exposures as determined by community response.
- 2 - * The range of exposures currently experienced during routine m i l i t a r y .
operations i s shown i n figure 1 9 , where t h e sonic-boom i n t e n s i t y i s indicated as a f'unction o f airplane a l t i t u d e f o r fighter and bomber aircraf't i n steady- l e v e l flight and i n maneuvers. Also shown are t h e estimated i n t e n s i t i e s for various proposed supersonic transports. It can be seen t h a t during routine m i l i t a r y maneuvers, exposures approaching 6 Ibs/sq f t have been experienced i n some communities. The proposed supersonic transports are designed on t h e basis of 2 lbs/sq f t during %ransonic acceleration and 1.5 lbs/sq f t during cruise. This requirement is based on very limited window br-age experience indicating t h a t t h e threshold of possible damage i s some- As pointed out i n the N A S A what g r e a t e r than 2 lbs/sq ft. (See ref.12 ) from these values S C m f e a s i b i l i t y studies, reduction of design overpressure specified w i l l r e s u l t i n severe range-payload penalties unless accomplished by b a s i c configuration improvements as previously discussed i n t h e section on Performance Aerodynamics (figure 6). It would be w e l l t o point out t h a t even though t h e estimated i n t e n s i t i e s f o r t h e proposed supersonic transports are w e l l within t h e range o f current exposures, t h e SST signatures w i l l have longer wave lengths. The importance of these longer wave lengths w i l l have t o be evaluated along with t h e wave form e f f e c t s discussed i n t h e succeeding paragraphs.
The community response aspect of t h e sonic-boom problem i s of prime importance and i s t h e most d i f f i c u l t t o evaluate becaus'e of i t s complex nature. For instance, people observe'not only the acoustic stimulus, but also are aware of sonic-boom induced vibrations i n building components and Their reactions are a function not only of t h e i r own observations, furnishings.
a l s o t h e environment i n which they live.
but * _- ~ .- - _ _ - - - _ _ ~- . .
- 3 - Important f-ndings of dhe recent FAA-Oklahoma C , , y t e s t s r e l a t e ;0 atmospheric e f f e c t s on overpressure values and wave form and t o building response. S t a t i s t i c a l r e s u l t s indicate t h a t the spread i n peak pressure presumably due t o turbulence and other atmospheric anomalies can be such t h a t about 15 percent of the t i m e , pressures may exceed mean values by about 30 percent of mean pressure and 1 percent of the time the pressures may exceed t h e mean values by 80 percent of t h e mean pressure. One of the characteristic b r e s u l t s of these tests i s thst the median values of peak pressure are The significance of the generally lower than the calculated nominalvalues.
pe& pressure is closely interrelated t o the pressure signature o r wave form.
t I n figure 20 are presented tracings of the measured waye forms from an accurately-calibrated and oriented array of matched microphones at separation distances of 200 feet. The wave forms are presented i n the proper time sequence and are d i r e c t l y comparable i n amplitude. These data illustrate t h e variations of wave forms obtained f o r given flights f o r which the a i r c r a f t operating conditions are essentially constant. It can be seen t h a t a wide . .: . d a distance on the ground of a f e w v a r i a t i o n i n wave shape occurred even over I hundred f e e t , and t h a t yariations were different f o r the two f l i g h t s . The peak overpressures value r i s e s and f a l l s as a function of distance i n much .- , . .
t h e same manner as the surface level of the ocean i n the presence of waves.
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Although not shown i n the figure, significant differences i n wave shape were measured at separation distances as small as 50 feet. Such variations as these, which have also been observed on other occasions (see ref. ll), are believed and velocity anomalies i n the atmosphere, p a r t i c u l a r l y t o result from temperature !
t h e lower layers. Invariably t h e highest measured pressures are associated.
* r I
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with highly-peaked wave forms, and conversely t h e lowest measured values are associated with rounded-off wave forms. It i s s i g n i f i c a n t t o note, however, that t h e impulse flmction y a r i e s over a much narrower range of values than does t h e corresponding peak pressure, and can exceed t h e mean value by 30 percent f o r about 1 percent of t h e time.
The overpressures of the magnitude encountered i n t h e Oklahoma City tests were not of s u f f i c i e n t magnitude t o cause primary s t r u c t u r a l damage Also, no t o w e l l constructed and w e l l maintained buildings (see ref. 1 3 ) .
appliances, or objects i n f o u r test damage occurred t o any f’urnishings, ho&s during t h e Oklahoma City t e s t s (ref. 13 ).
I . I I * SUMMAKY RENARKS I n summary, t h i s paper has reviewed a n u h e r of research areas i n support of t h e National Supersonic Transport Program and has indicated where improvements I n t h e area of performance aerodynamics i n the s t a t e of art appear attainable.
t h e o r e t i c a l methods f o r calculating and minimizing wave drag, drag-due-to-lift and sonic boom e f f e c t s have been programmed f o r use on high speed d i g i t a l computers. These methods and programs provide powerf'ul t o o l s with which it has been possible t o devise advanced aerodynamic configurations which exhibit f l i g h t efficiencies considerably higher than those previously demonstrated.
& I n the propulsion area mission and engine cycle studies have demonstrated t h e importance of careful airframe and propulsion system integration and -.. . * matching. There are large benefits t o be gained by improved component e f f i - ciency i n such areas as increased turbine i n l e t temperature, nozzle efficiency, etc.
S t a b i l i t y and control research indicates t h a t f l i g h t a t supersonic speed at high a l t i t u d e s presents problems i n proyiding adequate damping of t h e dynamic stability modes and probably w i l l require three axis damping augmentation. High of yaw-to-roll i n e r t i a r a t i o characteristic of the slender f'uselages values and mass distribution of t h e SST lead t o objectionable lateral-directional handling qualities. For fixed highly 5;zept arrow wings, t h e p o s s i b i l i t y of undesirable excursions i n angle-of -attack (pitch-up) can be minimized by u t i l i z a t i o n of t a i l o r e d wing leading-edge devices. Likewise f o r variable sweep configurations proper location of the horizontal t a i l i n combination with w i n g leading devices provides a meam t o minimize t h e problem.
I With regard t o structures and materials it has been determined t h a t .
design t o accommodate thermal stresses w i l l incur penalties of a f e w percent i n s t r u c t u r a l weight. Research on s t r u c t u r a l concepts indicates that t h e , l i g h t e s t structures can be obtained with skin-stringer construction u t i l i z i n g titanium alloys.
Ung time exposure a t temperature of 550% of titanium alloys and s t a i n l e s s s t e e l s exhibit no s i g n i f i c a n t damage i n mechanical properties. However, these tests have indicated some deterioration i n t h e .
t e n s i l e strength of spot welds.
Operating problem research is underway t o determine t h e compatibillty These studies have of t h e L . . SST with e x i s t i n g Air Traffic Control Systems.
used b o t h t r a f f i c control penetrations by a supersonic a i r c r a f t and SST ground based simulators flown by experienced a i r l i n e crews i n FAA controlled of t h e engine noise during take off simulated a i r t r a f f i c samples. Studies i n d i c a t e t h a t t h e leyels at the 3-mile point w i l l be comparable t o or l e s s than t h e subsofiic j e t s . For landing some compressor noise supression w i l l be required on t h e SST t o bring t h e noise l e v e l s below those of t h e current -I sulsonic j e t s .
Sonic boom research has indicated t h a t large penalties i n a i r c r a f t g r o s s : weight w i l l occur unless t h e SSTconfigurations are specially designed t o minimize t h i s effect. Research conducted i n t h e Oklahoma City sonic boom ?
program indicates t h e oyerpressures were not of s u f f i c i e n t magnitude t o cause pritnary structural damage t o well constructed and w e l l maintained buildings. Also, no damage occured t o any furnishings, appliances, o r objects i n t h e f o u r test houses.
; , 3 6 4 . . ' t
1 . Staff of the Langley Research Center: The Supersonic Transport -
A Technical Summary. NASA TN D-423, June 1 9 6 0 .
2 . Bisplinghoff, R. L . : The Supersonic Transport. Scientific American, June 1964, Volume 210, Number 6 .
3. Nichols, Mark R . : Supersonic Transport Propulsion Requirements.
Presented at the National Propulsion Meeting of the Institute of the Aero-Space Sciences,, Cleveland, Ohio. March 9-10, 1 9 6 1 .
4 . Nichols, Mark R . : The Supersonic Transport - Required Characteristics
of Configurations. Society of Automative Engineers. Paper 341F 1961.
SAE National Aeronautic Meeting. New York, New York.
5 . White, Maurice D., Sadoff, Melvin, Bray, Richard S . , and Cooper, George E . : Assessment of Critical Problem Areas of the Supersonic Transport by Means of Piloted Simulation. Aerospace Engineering, Vol. 21, No. 5, May 1 9 6 2 .
-.. ..
6 .
Clark, D . C . , Notess, C . B., Pritchard, F . E . , Reynolds, P . A,, and Schuler, J . M . : Application of Self-Adaptive Control Techniques to the Flexible Supersonic Transport. ASD-TDR-63-331, Vol. 1, August 1963.
7 . Spearmh, M . L . , and Driver, Cornelius: Some Factors Affecting the
Stability and Performance Characteristics of Canard Aircraft Configurations. NACA F M U8D16. June 1 9 5 8 .
8. Baals, Donald D . , and Polhamus, E. C . : Variable Sweep Aircraft.
Astronautics and Aerospace Engineering, June 1963.
Sawyer, Richard H . , Stickle, Joseph W., and Morris, Richard: A Simulator 9.
Study .of the Supersonic Transport in. the Air Traffic Control System.
Proceedings of the National Aerospace Electronics Conference.
Vmy 11-13, 1 9 6 4 . Dayton, Ohio.
1 0 .
Trapp, W . J., and Forney, D . M . , Jr., eds: WADC-University of Minnesota Conference on Acoustical Fatigue. WADC Tech. Rep. 59-676, U . S . Air Force, March 1 9 6 1 .
1 1 . Hubbard, Harvey H . , Maylieri, Domenic J., Euckel, Vera, and Hilton, David A . : Ground Measurements of Sonic-Boom Pressures for the Altitude NASA TR R-198, July 1 9 6 4 .
Range of 1 0 , 0 0 0 to' 75,000 feet.
12. Power, J . K . , and Bates, George: Sonic Boom and Community Relations.
(preprint) 689, SOC. Automative Eng., April 1963.
13. Power, J . K . : Sonic Boom Effects on Light Aircraft, Helicopters and Ground Structures. FAA Office of Supersonic-Transport Development, Washington, D . C . (paper presented at American Society for Testing and Materials, June 25, 1 9 6 4 ) .
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