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NASA Technical Paper 1805

of.anA.spect-Ratio-10Supercritical-.'Wing Transport' Model Equipped .. .

. - ith a. Full-Span Slat and Part-Span-1 and-Full-Span .Double-Slotted,Flaps

Harry L.. Morgan, Jr. . . ,

APRIL 1981 FOREARLY DOMESTICDISSEMINATION

Nnsn

Because of ts significant early commercial potential,,.dhisinformation, which has been developed under a U.S. Gdv-ernment program, is being disseminated within the Ud tedStates in advance of general publication. This inform&"ionmay be duplicated and used by the recipient with the ex-press imitat ion hat t not be pu-blished. Release of this. . .

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. . :.' . Review for generalelease March 31, 1 9 8 3 ,


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FEDD DOCUMENT

Note that this document b e a r s the abel l lFEDD, l l anacronym for FOR E A R L Y DOMESTIC DISSEMINATION.The FEDD l a b e l is a f f i xe d todocuments hat may containinformation having high commercialpotential.The FEDD concept w a s developed a s a result o f thedesire to maintain U.S. leadership in world rademarketsandencourage a favorable.balance o f trade. Since theavailability o f tax-supported U.S, technology o oreignbusiness nterestscould epresent anunearnedbenefit,research results that may have high commercialpotentialare being distributed to ti.S. industry i n advance of generalre lease .The recipient o f this reportmust reat he informationit containsaccording to theconditions o f the FEDD labelon the front cover .


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TECH LIBRARY KAFB,NU

NASA Technical Paper 1805

Low-Speed AerodynamicPerformanceof an Aspect-Ratio-10Supercritical-Wing Transport Model EquippedWith aFull-Span Slat and Part-SpanandFull-Span Double-Slotted Flaps

Harry L. Morgan, Jr.Langley ResearchCenterHamnptol-1 Virginia

National Aeronauticsand Space AdministrationScientific andTechnicalInformation Branch1981


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SUMMARYAn in ve st ig at io n was conducted n heLangley 4- by7-Meter Tunnel t ode te rmine he s t a t i c longi tud ina l and la te ra l -d i rec t iona laerodynamiccharac-t e r i s t i c s of an advanced aspect- ra t io-10 supercr i t ical-wing ranspor t modelequippedwith a full-span eading-edge s l a t as well as part-spanandfull-spant r a i l ing - edge l aps .T h i swidebody r anspor t model was also equippedwithspoiler and a i l e r on r o l l - con t r o l su r f aces, f law-through nacelles , andinggear ,andmovable horizontal t a i l s . S i x ba s i c wing conf i gura t ions were te s teddur -ing h i s nve s t i ga t ion and cons i s tedof 1)c ru ise ( s l a t s and f lapsn e s t e d ) ,(2)cl imb ( s l a t s def lec tedand f l ap s nes ted ) , ( 3 ) par t - span lap , ( 4 ) fu l l - spanf l a p , (5) f u l l - s p a n l a pwi th law-speed a i le ro ns , and 6) ul l- span lapwi thhigh-speedailerons. Each of the our lappe d wing conf i gura t ions was tested

with he eading-edge s l a t and the ra i l ing -edge f lapsde f lec ted t o s e t t i n g sr e p r e s e n t a t i v e of both ake-offand andingconditions. The tests were con-ducted a t fyee-s treamcondi t ionscorresponding t o Reyn olds numbers (basedont h e mean geom etri cchord)of 0.97 t o 1.63 x l o 6 and corresponding Mach numbersof 0.12 t o 0.20, hroughanangle-of-attack angeof.-4O t o 24O and a s i d e s l i pangle angeof - l o o to 5O. The par t - and ull-span wing con f ig ura tio ns werea l so t es ted i n groundproximity.The longi tud ina l t e s t r e s u l t s show t h a t a l l th e wing conf igura t ion s t e s tedexhib i ted wing- t ip s t a l l beha vior oll cwed by a r e d u c t i o n n o n g i t u d i n a l sta-b i l i ty . With e i the r ake - o f f or l a n d i n g l a pse t t ings and a t a given untrimmed

lift coe f f i c ien t , he h r eef u l l - span f lapconf igu r a t ions p r oduced more nega-t i ve pitching-moment coe f f i c ien t s han he pa r t - span f l ap con f i gu r a t ions and ,there fore , ncur redhigher trim dr agpena l t i e s . A comparativeanalysisof hetrimmed pe r f o r mancecha r ac te r i s t i c s of the f ou rf lapp ed wing conf i gura t ionstes ted i n d i c a t e s h a t h ec o n f i g u r a t i o n w i t hf u l l - s p a n f laps and low-speeda i l e r o n s had s l i g h t l y bet te r trimmed per formance han heother hreef lappedconf igurations . The l a t e r a l t e s t r e s u l t s show t h a t h e a t e r a l - d i r e c t i o n a ls t a b i l i t y o feachf lapped wing con f i gu r a t ion wi th l and i ng f l ap s e t t i ngs wass l i g h t l y less than hes tab i l i ty o f heco r r e spond ing conf igu r a t ion w i t ht a k e - o f f l a ps e t t i n g s . For the u l l - s pan l ap wing con f igu r a t ionw i t he i the rtake-off or l a n d i n g f l a p s e t t i n g s , a r g e d e f l e c t i o n s o f h e e f t o u t b o a r d ro l l -c o n t r o l spoilers producedchanges i n roll ing-momentcoefficient as g r e a t asthosep roduced by d i f f e r en t i a l de f l ec t i on s o f he law-speed a i l e r o n s f o r h ecor respondingpar t- span lap wing conf iguratio ns .Largedef lectionsof her o l l - c o n t r o l spoilers a l so r e s u l t e d i n anunfavorable loss o f l i f t , a p o s i t i v esh i f t i n p i t c h i ng moment, andan ncrease nnega t iveyawing moment.

INTRODUCTIONThe rapid worldwide in c r ea se in th e consumptionand pr ice ofcrude o i l i nrecent yearshasgenera ted a renewed i n t e r e s t bymany governmentandprivater e sea r c h o r gan i za t io ns in ways of improving he energy e f f ic iency of vehic lest h a t u s e f u e l s d i s t i l l e d f r o mcrude o i l . I np a r t i c u l a r , NASA hasbeenac t ive ly


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i nvo lved nanae ronau t i ca l research project to improve t he ene rgye f f i c i encyof modern wide-body j e t t r a n s p o r t a i rc ra f t . The Aircraft EnergyEfficiency(ACEE) project was formulated to stimulate research efforts by both i ndus t ryand NASA. One element of t h e ACEE project is t he EnergyEf f i c i en tTranspor t(EET) program which is concernedpr imar i ly wi th t h e development of advancedaerodynamic and ac t ive-cont ro ls echnology for a p p l i c a t i o n to d e r i v a t i v e ornex t -gene ra t ion r anspor t aircraft . One par t of the EET program hasbeen t h eaerodynamicdevelopment, by NASA LangleyResearchCenter (LaRC) personnel , ofadvanced su per cr i t i ca l wi ngs wi t h greater sec t ion h ickness-chord ra t ios , higheraspect ratios, h i g h e rc r u i s e l i f t c o e f f i c i e n t s , and lower sweepback t han theconvent iona lwingsoncurrent ranspor ts . These advanced superc r i t i ca l wingshavebeen tested e x t e n s i v e l y n t h e LaRC wind tunnels to determine their high-speed cruise performance characterist ics (refs. 1 and 2 ) . Because of t h e i rh i g hcruise l i f t c o e f f i c i e n t s andhigh aspect ratios, these wingscould be smallerand mre f u e l e f f i c i en t hanconven t iona lwings ,p rov ided t h e low-speed, high-lift performance requirements could be met.

To determine he law-speed performance characteristics of a r e p r e s e n t a t i v ehigh-aspec t - ra t io supercritical wing, a 3.66-m (12-ft)span law-speed j e t t r ans -port model was fabr ica ted and tes ted in the Langley 4- by7-Meter Tunnel asreported i n r e f e r e n c e 3. T h i s model was equipped w i t h a conven t iona l lys i zedpart-span, double-slotted t r a i l i ng -edge l apsys t em, ful l-span leading-edge s l a tlow- and high-speedai le rons , spoilers, and in te rchangeable aspect-ratio-12 andaspect-ratio-10 wing t i p s . The p r e s e n t n v e s t i g a t i o n was conducted t o determinet h e low-speed performance characterist ics of t he a spec t- r a t io -10vers ion oft h i s model equipped w i t h both a part-spanandful l -span, double-slotted t r a i l i n g -edgeflapsystem. The model tested was a 3.23-m (10.59-f t span model of anadvanced ong-range, wide-body j e t t r a n s p o r t wi t h cru ise wingand fuselagedimensions similar to those of t h e NASA SCW-2c s u p e r c r i t i c a l wing model testedin the Langley8-FootTransonic Pressure Tunnelandrepor ted nreferences 1and 2. T h i s wing had an aspect r a t io of 10, a 27O qu art er -c ho rd sweep, andstreamwise supercri t ical a i r f o i l s e c t i o n s t h a t v a r i e d n maximum thi ckn ess -ch ordratio from approximately0.15 a t t h e wing root to 0.107 a t t h e wing t i p .

The basic h i g h - l i f t f l a p sys temcons is ted of both part-spanand ful l-spandouble-slotted t r a i l i ng -edge f l aps and a ful l -span eading-edge s l a t . Thet r a i l i ng -edge f l a p c o n s i s t e d of anadvanceddesign large vaneand small a f t f l a pcombination, as opposed to t h e more conventionalcombinat ion of small vaneandl a r g e l a p . Also as p a r t of the EET program, a s imi l a r lydes igned a rgevaneand small a f t f l a p combination was tested e x t e n s i v e l y by t h e Douglas Aircraf tCompany and is reported i n e f e r e n c e 4 . The par t - span f l a p conf igu ra t ion wasa l so equipped w i t h inboard high-speedai le rons , outboard low-speed a i l e r o n s ,and spoi lers a t the inboardand outboard f l ap l o c a t i o n s . The fu l l - span f l a pconf igu ra t ion was obta ine d by replac ing t he high- and lm-speed a i le ronsegmentsw i t h double-slotted f l a p segmentsequipped wi t h spoilers. Both part-and f u l l -span f l a p c o n f i g u r a t i o n s were also equipped w i t h two wing-mounted flaw-throughn a c e l l e s , a n d i n gg e a r ,f i x e dv e r t i c a l t a i l , andmovable horizontal t a i l s .

Two a d d i t i o n a lh i g h - l i f t f l a p c o n f i g u r a t i o n s were tested during t h i s inves-t i g a t i o n . The f i r s t was obt aine d by rep laci ng he nbo ardhigh-speedai le ronsegment of the part-span f l a p system wi t h a double-slotted f l a p segmentand wasdes igna ted as t h e fu l l - span f l a p with low-speed a i le ro ns wing conf igu ra t ion .2


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The second was obta ined by repl acing he outb oard law-speed aileron segment ofthe par t-span f l a p system w i t h a double-s lottedflapsegmentand was des igna tedas the u l l - span f l a p withhigh-speedaile rons wing conf i gurat ion. Each of hef o u r flapped wing conf igurations was tested w i t h h e f l a p elements s e t a t amoderate d e f l e c t i o n t o represent ake-offcondi t ions and a t a h igh de f lec t ionto represent landingcondit ions . Eachof the our lapp ed wing conf i gura t ionswas tes ted with hefull-span eading-edge s l a t f u l ly de f lec ted ; hepa r t - spanand fu l l - span f lap wing conf ig ura t i ons were also tested w i t h t h e s l a t nested.I n a d d i t i o n t o the four f lapp ed wing conf igur a t io ns , a c r u i s e wing con f igu r a t ion( s l a t s and flapsnes ted)and a climb wing conf iguration (slats f u l l y deflectedand f l a ps nes ted) were a lso tested. A t o t a l of 10 wing conf igura t ions weretested: (1) c r u i s e , (2) climb, ( 3 , 4) pa r t - span l ap take-off andanding,(5, 6 ) f u l l - span l ap ake - o f f and landing , (7, 8) f u l l - s p a n l a pwith law-speeda i l e r o n s t a k e o f f and andi ng, and (9, 10) fu l l - span f l a p withhigh-speed a i l e -rons ake-offand anding.

The cruise, climb, part-span take-of f , and par t-span landing f l a p wingcon-f i g u r a t i o n s were also t e s t ed dur ing a pr ev ious nves t iga t ion reported i n r e f e r -ence 3. As p o i n t e do u t n h ed i s c u s s i o ns e c t i o n of t h a t report, a thoroughcheck of thevaneand f l a p p o s i t i o n i n g a f t e r c o m p l e t i o n o f t h e tunne l testsrevealedan error of 0 .6 an (0.25 in . ) i n t h e l a t e r a l d isp lacemen t of t h e l e f twing in boa rd lapsys tem.Th is error i n l a t e r a l displacement was correctedp r i o r t o t h i s n v e s t i g a t i o n , and selected tests of hepar t - spanflap wingcon-f i g u r a t i o n w i t h both ake-offand anding f l a p s e t t i n g s were repeated.T h i s i n v e s t i g a t i o n was conduc ted n t h e Langley 4- by 7-Meter Tunne l a tf ree-s treamcondi t ionscorresponding t o Reynoldsnumbers (based on he meangeometricchord)of 0.97 to 1.63 x 106 and co rr es po nd in g Mach numbers of 0.12t o 0.20, throughanangle-of-attack angeof -4O t o 24O and a s i d e s l i p a n g l erangeof -1 Oo to So. The part- and f u l l - spa n , ake of f and anding f l a p wingconf igu r a t ions were also tested in p r ox imi ty o f he unne l f loor t o s i m u l a t egroundeffects. The model was ins t rumentedwith a six-component train-gagebalance to measure aerodynamicforcesan d moments and w i t h chordwise s u r f a c es tat ic-pressure t a p s a t th r ee spanwise s ta t ions to de te rmine representa t ivewing and f l a p l oads . The pressure data o b t a i n e dd u r i n g h i s n v e s t i g a t i o n arep r e s e n t e d ng r aph i c and t abu l a r form i n e f e r e n c e 5. This report presentsandd iscus se s he s t a t i c long i tud ina l and l a te r a l - d i r ec t iona laerodynamic da t ao b t a i n e d d u r i n g h i s n v e s t i g a t i o n .

SYMBOLS AND ABBREVIATIONSThe lon gi t ud i na l for ces and moments p res en te d in t h i s report are re fe rencedto the s t ab i l i ty - ax i s sys te m and the l a t e r a l forces and moments t o t h e body-axissystem.The moment data are referred to a moment center located on he modelc e n t e r i n e ( i n t e r s e c t i o n o f h e wing r e f e r en c eplane and model symmetry plane)a t t h e 1.64-m (5.39-ft) body s ta t ion, which is 46.72 cm (18.39 in . ) ongi tud in-a l l y a f t of the wing root lead ing edge. The lon gi t ud i na l oca t io no f h e momentcenter corresponds t o the qua r te r - cho r d po in t oca t ion of t h e mean ge omet ricchordof he rapezoida l wingplanform planformwithout ra i l ing-edgeextension)whichextends f ran t h e model c e n t e r i n e t o the wing t i p . The aerodynamiccoef-f i c i e n t data are based on the t rape zoid a l wingplanformwhichhas a r e f e r ence

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area of 1.04 m 2 (11.21 f t 2 ) , a referencespan of 3.23 m (10.59 f t ) , and are fe re nc e mean geometric chord of 34.1 4 cm (1 3.44 in . ) .A l l measurementsandcalculations were made i n U.S. CustomaryUnits.V a l u e s presentedherein are g i v e n n t h e In te rna t iona l Sys temofUni ts ( S I ) ,

with t h e equ iva len t va lue s n U . S . Customary Uni tsgivenparen the t ica l ly .

c

Cn

b2aspect ra t io , -Swing span, m (ft)l o c a l streamwise wing chord, can ( in . )r e f e r ence mean g em e t r i c chord, cm ( in . )

Dragqsd r a g o e f f i c i e n t , C D in omputer-generat ed t ab les )

L i f tq sl i f t c o e ff i ci en t , CL in omputer-generat edables)

Rolling momentro l l ing-momentoeff ic ient , (CRM inomputer-generatedqSbt ab les )e f f e c t i v ed i h e d r a l parameter based on ncrement of C1 betweenac 2

aB6 = -loo and 5O; 0 I 1/degPitching momentpitching-momentoeff ic ient , (CPM inomputer-generatedqsc

t ab les )Yawing momentyawing-moment coefficient, (CYM inomputer-generatedqSb

t a b l e s )d i r e c t i o n a l t a b i l i t yp a r a m e t e r based on ncrement of C n betweenac,6 = - loo and So; 9 I I/deg..a6

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SideforceCY

cyB

e

it

B

s ide-forceoefficient , (CSF i n computer-generated tables)qsside-force parameter based on increment of Cy between B = -loo

acya Bnd 5O; 0 , 1/deg

lon git ud ina l sta bil ity parameter (CMCL i n computer-generated tables)

CL2wing ef f i c iency ac to r , CD = C D , ~ -, where CD,o is drag coef-TAef i c i en t a t zero l i f t

height-to-span r a t i o ofmodelmoment ref erence cente r above f lo o r,m ( f t ) H/B i n computer-generated tables)

incidence of ho riz on tal ail ,posi t ivefo r ea di ng edgeup,deg(ISUBT i n computer-generated tables)

l i f t -d rag ra t i o (L/D i n computer-generated tables)free-stream Machnumber (MACH i n computer-generated tables)free-stream dynamic pressu re , kPa ( l b / f t 2 ) (Q i n computer-generatedtab les )free -s tream Reynolds number based on cwing referencearea, m 2 ( f t 2 )

-

wing thickness-chord ratioangle of a tt ac k ofmodel referencecenter ine,positive noseup,deg(ALPHA i n computer-generated tables)angle of s id es li p ofmodel referencecenter ine,posi t ive nose l e f t ,deg BETA i n computer-generated tables)ai lerondeflect ionangle ,posi t ive fo r ra ili ng edgedown,degflap def lec t ion angle , pos i t ive for t ra i l ing edge down, degsla t de fle cti on an gle , po sit ive fo r ra ilin g edge down,degspoiler deflect ion angle, posi t ive for rai l ing edgeup,deg

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6, va nee f lec t ionng le ,o s i t ive for t r a i l i n g edge down, degrl nondimensional wing semispanocationSubscripts:corr correctedR l e f tmaxK r i g h tAbbreviat ions:L.E. leadingdgeT.E. t r a i l i n gd g eW.R.P. wing referencelane

MODEL DESCRIPTIONThe h i g h - l i f t research model tested dur ing t h i s i n v e s t i g a t i o n had a 3.23-m(10.59-ft)spanand was r e p r e s e n t a t i v e of anadvanced ong- range, wide-bodyj e t t r a n s p o r t wi t h cruise wingand fuselage dimensions scaled from those of t h e

NASA SCW-2c high-aspect-ratio superc r i t i ca l model developed a t t h e NASA LangleyResearch Centerand reported i n e f e r e n c e 1. The wing was fabr icated w i t hremovable eading-and railing-edge egments. The c ru i s e wing egments couldbe removed easily and replaced w i t h a leading-edge s l a t and ra i l ing-edgespoi ler/ f lap and aileronsegments.Alt houg h many wing co nf ig ur at io ns were pos-s ib le , s i x basic wing conf igurations were tested dur ing t h i s inves t iga t ion :(1 ) cruise ( s la t s and f lapsn e s t e d ) , ( 2 ) climb (slats deflected and f lapsnested) , 3)par t-span f l a ps , ( 4 ) fu l l - span f l aps , (5) fu l l - span f laps wi t h low-speed a i le rons , and ( 6 ) fu l l - span f laps wi t h high-speed a i l e r o n s . Each of t h efour flapped wingconf igura t ions was tested wi t h t h e f ull-span leading-edge s l a tand the t ra i l ing-edgef lapsegments deflected to s e t t i n g s r e p r e s e n t a t i v e ofb o t h take-off and andingcondit ions . A deta i led wingplanform ayout of t h ebasic control and f l a p surfaces is p r e s e n t e d n f i g u r e l ( a ) ; a s k e t c h of t h e s i xbasic wing conf igura t ions tested i s p r e s e n t e d n i g u r e l ( b ) . Photographs of t hemodel i n s t a l l e d n t h e Langley 4- by 7-Meter Tunnel are shown i n f i gu r e 2. Thep e r t i n e n t model geometric characteristics are summarized i n table I. Detailedwingand f l a p component surface coord ina tes are g i v e n n r e f e r e n c e s 5 and 6 .

The model was fabr icated w i t h aluminum wings, a g lass f iber fuselage,andan empennage for minimaldef lections a t the de s igncondi t ions of a maximumtunneldynamic pressure of 2.87 kPa (60.0 l b / f t 2 ) and a maximum wing l i f t coef-f i c i e n t of 3.0.The mpennage co n si st ed of movable horizontal t a i l s withoute leva tors and a f i x e d v e r t i c a l f i n without a rudder. The hor izontal t a i l s weremountedon t h e model w i t h a geared,p ivot ing bracket t h a t allowed for incidence6


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ang le s from -15O to 1 5 O i n 5 O increments. The model was also equippedwith twowing-mounted, flow-throughnacelleswi th scaled externald imens ions similar tothose of a typical high-bypass r a t i o (approximately 6) turbofanengineusedoncu rr en t wide-body je t s . The model was a lso equippedwithsimulated andinggear and doors a t tached to the wingand fuselageundersidenear henose.The bas ic cruise wing was designedwith an aspec t- ra t io -10 rapezoida lplanformwhichextended rom he model c e n t e r l i n e to th e wing t i p andhad27O quarter-chord sweep. The winghadan nboard ra i l ing-edgeextension hats ta r ted a t t h e ll = 0.434wing semi span tati onand ncreased hechord a t t h ece nt er in e by 40 perce nt . Thewinghad streamwise s u p e r c r i t i c a ls e c t i o n sw i t hmaximum thickness-chord ra t ios of 0.144 a t theside-of-bodysemispan ocation

(11 = 0.109) , 0.120 a t the r a i l ing - edge b reak s t a t io n (Tl = 0.434) , nd0.107a t the wing t i p (17 = 1.0). Thewing was mounted on he us el ag ew i t h a 5 O dihe-d ra l angle and a - lo inc idenceangle a t the wing cente r ine .Cont ro l andFlapSystems

The leading-edge s l a t , t r a i l ing-edge f l ap , and spoiler and a i le ronc o n t r o ls u r f a c e areas were s ized and posit ionedspanwiseon he ba sis of a comparativea n a l y s i s o fs e v e r a le x i s t i n g d e s i g n s for lower aspect- ra t io-6 to aspect-ratio-8transpor twings . The t r a i l i n g - e d g e l a p had a doub l e - s lo t t ed l ap ha t con-sisted of anadvanceddesign argevane and small a f t f l a p combina t ion n c o m -pa r i son wi th he more convent iona l small vaneand la rge a f t f lapcombina t ions .Advanced designs similar to t h i s combinationhave ecentlybeenunderdevelop-ment by severala i r c r a f tm a n u f a c t u r e r s ( r e f . 4 ) andhaveexperimentallyachievedmaximum two-dimensional l i f t coef f ic ien ts approaching hoseachieved by th e morecomplex t r ip le - s lo t te d laps y s t e m s . The s t r u c t u r a l loads produced by t h i s f l a pcombination are less severe han hoseof heconvent iona l combina t ions becausea gr ea te r pe r cen tage o f he t o t a l vane/f lap loads are generated by th e morec lose lycoupled a rgevane component.

For t h ep a r t - s p a n f l a pc o n f i g u r a t i o n , a s imple-hinged,high-speedaileronsegment was pos i t ioned outboard of the break s ta t ion , and a simple-hinged, low-speedaileronsegment was pos i t ionedoutboard o f the ou tb oa rd f l ap segment.The l e f t and righth igh-speeda i le ronscould be deflected from -30 t o 50, andt h e e f t and r i g h t low-speed a i le ro ns , from -30 to 30. Both l e f t a n d r i g h tinboard lapsegments were equippedwithground spoilers , and both l e f t andr i g h to u t b o a r d f l a p segments were equippedwi th l igh t spoilers . The l e f t andr igh t g r o u n d and f l i g h t spoilers could be deflected to e i t h e r 45O or 60, whichare pr imar i lyground if t- lossand speed-break de f le c t ion s . The l e f to u t b o a r df l i g h t spoiler could a lso be def lectedf rom Oo to 20 i n 4O increment s,whichare p r i m a r i l y f l i g h t r o l l - c o n t r o l d e f l e c t i o n s .

To ob ta in he h r eef u l l - span f l a p conf igu r a t ions tes ted, t h e l e f t and r igh thigh-and/or ow-speedaileronsegmentsof hepart-span f l a p conf igu r a t ion werereplaced with proper ly contoured double - s lo t ted f l a p segmentsequippedwithf l i g h t spoilers. Both l e f t and r i g h t , f l i g h t spoilers could be def lected e i t h e r45O or 60, and the l e f t f l i g h t spoilers could a l so be d e f l e c t e d from Oo to 20i n 4O increments . The f l i g h t spoilers of t h e f l a p segmentused to replace t h el e f t low-speed aileron segment w i l l be called " r o l l - con t r o l spoilers.

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Slat and Flap SettingsThe s l a t , vane, and f l a p componentswere set at def lect i ons rep rese ntat iveof e it he r climb, take -off , or landing wing configurations, and the correspondinggaps and overlaps were then optimized for maximum l i f t using he heoret ical

two-dimensional,multicomponent a ir fo il an a ly s is program described i n re fe r -ence 7 . . A ketch of the de fle cti on , gap, and ov er lap de fin iti on employed duringt h i s investigation is presented i n f igure 3. The def le ct io ns and ov erla psaredefined rel ati ve o he on ge st chord of the pa rt ic u la r components. The lon gestchord is defined as he dis tan ce from the midpoint of the ra ilin g-e dg e base ofthe component to he. forward-most leading-edge co ordina te . The ov er lap isdefined as he dis tan ce from the lower surf ace railin g-ed ge coo rdin ate of theforward component along i t s long est chord to a po int a t which a perpen diculardropped from th a t chord in te rs ec ts th e forward-most co ord ina te on the ead ingedgeof the a f t component.The gap is defined as the shortestdi st an ce from thelower surface railin g-ed geco ordina te of the forward component to the uppersu rf ac e of the a f t component.

Thecomponent geom etries of theflapped wing sec tion s at the trailin g-e dgebreak station (r l = 0.434) were used to perform th e he or e ti cal two-dimensionalgap and overlapoptimiz ations. The re su lt s of theoptimizationsare summarizedi n the ollowing tab le:

wing Deflection,configuration Component deg W?/cC l i m bTake-of f

Landing

S l a t 0.0250S l a t -50

.Ol5lap

.0155ane 0.02

S l a t -50 0 . 0 2Vane 30 .02Flap .Ol0

Over lap/c0.020.02.04.Ol

0.02.03.005

These wo-dimensional def le ct io ns, gaps, and ov erlap s were incorporated ntoth e ac tu al three-dimensional wing using positioning igs ocated a t the edgesof the leading-edge s l a t and tra ili ng -edg e lap segments. For each wing con-figu ratio n, he railin g-ed ge vaneand f la p component def le ct io ns , gaps, andoverlaps a t the railing-edge break sta tio n were s e t in i t i a l l y , and thedeflec-tio ns of thepositioning j i g s for he inboard and outboard segments were thenindividually adjusted i n the streamwise di re ct io n to main tain he proper gapsand ove rlap s. The change i n de fl ect io ns or he inboard and outboard segmentswas necessary because of the geometric twist of th e wing.The trailing-edgevaneand f l ap componentswere als o tra ns lat ed s l i g h t l y i n an attempt to keeptheir pressure aps i n the same streamwise plane a s th a t of the mainwing taps.The gaps and ove rlaps of the inboard vaneand f l a p componentswere s e t a tconstan t valu es along th e segmentspanandwerebasedon the local cruise wing8


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chord a t the ra i l ing-edge break s t a t i o n . The gapsandover laps of theoutboardvaneand f l a p components were set a t cons tan t percentage va lues based on theloca l chord.The trailing-edgevane /f lap brackets were a t t a c h & n p l a n e s p a ra l l e l to

t h e wing ymmetry pl ane . The leading-edge s l a t was p o s i t i o n e d n a similar man-nerexcep t hat he nboar d and outboard segments were adjus ted i n p l a n e s per-pendicu la r to the eading-edgeof t h e cr ui se wing.The gapandoverlap werese t a t cons tan tpercentage va lues based on he iocal chordof he wing wit hou tthe ra i l ing-edgeextens ion t rapezoida lpla nfo rm) . The leading-edge s l a tbrackets were a t t a ched np lanespe r pend icu la r to the wing leadingedge. Noa t t empts were made to account for t h e d e f l e c t i o n o f e i t h e r s l a t or f l a p bracketsunderaerodynamic oading.

Wing Pressure TapsThe pressure data o b t a i n e d d u r i n g t h i s i n v e s t i g a t i o n a re pr e sen ted n bo thgraph ic and tabu lar orm n eferen ce 5. As i l l u s t r a t e d n i g u r e l ( b ) , t hewing was ins trumented w i t h chordwise rows of surface s t a t i c - p r e s s u r e taps a tth reespanwises ta t ions labe led A, B, and C and located a t rl = 0.266,0.624,and0.907, respectively.S ta t ions A and B each had 70 pr ess ure aps or bo thpar t - and fu l l - s pan lap wing con f ig ura t io ns .S t a t i o n C had 47 taps for t h epar t - spanfla p wing confi gurat ionand 49 f o r t h e f u l l - s p a n fl ap wing config-ura t ion .Several componentcombinations were possible a t eachof t h e th rees ta t ions and are i l l u s t r a t e d n f igu r e 4 . For t h e par t-s pan lap wing conf ig-u r a t i o n s , a l l the combina t ions p r e sen ted nf igu r e 4 were possible a t sta-t i o n s A and B; however,onlycombinationsusingcomponents A, E, and F were

possible a t s t a t i o n C. All component combinationsp r e sen ted n igu r e 4 werep o s s i b l e a t a l l th re e s t a t i o n s for thef u l l - span f l a p wing conf igurations .

TESTS AND CORRECTIONSThe tests were conducted i n t h e Langley 4- by 7-Meter Tunnel,which hasa test s e c t i o n of 4.42 m (14.50 f t ) by 6.63 m (21.75 f t ) . These tests wereconducted a t free-streamdynamicpress ures rom 0.96 to 2.87 kPa (20.0 t o60.0 l b / f t 2 ) . CorrespondingReynoldsnumbers, based on th e ef er en ce meangeometricchordof 34.14 c m (13.44 in.) , were 0.97 t o 1.63 x 106and corres-ponding Mach numbers were 0.12 t o 0.20.The model was tested throughanangle-of-attack range of -4O to 24O and a s ides l ip -angle range of - loo to So.The aero dyna mic orcesand moments were measured by a six-componentstrain-gagebalance mounted i ns id e t h e fuselage. The angleof a t t a c k was se t by thep i t c h d r i v e of the model support systemandmeasured by an e l ec t r on ic in c l i -nometermounted i ns id e he forward por t ionof he use lag e. The s ide s l ip anglewas s e t by the yaw drive of the modelsupportsystemand was measuredbyane lec t r on iccounter mountedon t h e yaw-drivegearing ystem. Thewing surfaces t a t i c pr e s su r e s were measured by e i the r 17.24- or 34.47-kPa (2.5 or 5.0 lb/in2)d i f f e r e n t i a lp r e s su r e r ans duc e r s and s ix48-por t canningvalves .Fuselagechamber and base pr e s su r e s were measured by 6.89-kPa (1 .O l b / i n 2 ) d i f f e r e n t i a lpressure t ransducers .

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Eoundary-layer transition strips were located 2.54 cm (1.0 in.) normal tothe leading edge n both upper and lower urfaces of the cruise ing configu-ration only, n the horizontal and vertical tails, nd on the outer surfacesof the flaw-through nacelles. The transition roughness was sized accordingoreference 8 and required a commercial No. 60 abrasive grit sparsely applied.Wind-tunnel jet-boundary corrections were computed according o refer-ences 9, 10, and 1 1 , and the averaged values re applied to the force and momentdata. The corrections were applied as follows:

cD,corr = cD + JlcL2Cm,cOrr = C + J3CL (for tail-on data)

where J1 = 0.0045, J2 = 0.2581, and J3 = 0.011. Wing, body, and wake solidblockage corrections were also applied to the datand were determined accordinto reference 12. Drag corrections due to model chamber and base pressures refer-enced to free-stream static pressure were also applied o the data. No correc-tions for tunnel flaw angularity were made to the data because no provisionswere made to test the odel in the inverted position. However, flow angularitymeasurements, made. uring several revious investigations on similar models posi-tioned at the same approximate locationn the tunnel, showed that a. l o to 0.2Oup-flow correction was equired.

PRESENTATION OF RESULTSAlthough numerous test variablesnd wing configurations were possible forthis high-lift research model, only combinations representativef the more sig-nificant configurations were tested during this investigation. Theest resultsfrom a prior investigation as presented in reference 3 show the effects of

(1) an increase in aspect ratio to 12, (2) nacelles on/off, (3) transitionstrips on/off, (4) spoiler deflection on longitudinal lift loss, and (5) high-speed aileron deflection on lateral-directionalharacteristics for the cruise,climb, and part-span flap wing configurations. The particular longitudinal andlateral-directional test variables and wing-configuration combinations testedduring this investigation re presented in the following table:

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Figure index for wing configurations of -:mise T T T

~ ~ ~~~

Full-span flap Full-span fl ap withlov-speed aileron

Full-span fla p wit hhigh-speed aileronart-span flapext variable

Longitudinaldata:Reynolds nmber

( S u a n a q , fig. 11)Eorizontal-tail deflection

SLIIIarLTy, fig. 18)Landing gear on/offSlat deflection/nestedGround effects

Lateral data:Sideslip angle-speed aile ron &fl ecti on(roll control)Spoiler &flection(roll control)

ILandingakeoff Landing Takeoff Landing Takeoff

9 (a)

16(a), (b)

19(e).

26(a), (b)

6

13

2 6 ( C ) , (d:

""""_.

23 24(a) , (b)8(a) o d)

. _


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Unlessotherwise s t a t ed on t h e data f i g u r e , t h e n a c e l l e s were on,and t h eh o r i z o n t a l t a i l s were off for a l l wing conf igura t ions tested. I na d d i t i o n ,unlessotherwise s ta ted , thegear was of f for t he c r u i s ea n d climb wing config-ura t ions and was on for t h e f lapped wing configurat ions. L i s t e d on each f i g u r eare t h e runnumberscorresponding to t h e data plot ted. The t abu la t ed ong i -t u d i n a l s t a b i l i t y - a x i s a n d l a t e r a l body-axis d a t a for a l l t h e r u n sp r e s e n t e d nt h i s reportare given nappendix A. The trim l o n g i t u d i n a ls t a b i l i t y - a x i sd a t aobta ined by in t e rp ola t io n of t h e tes t d a t a for v a r i o u sh o r i z o n t a l - t a i l deflec-t i o n s are g iven n append ix B.

DISCUSSION OF RESULTSThe d i scuss ionof t h e t e s t resul ts i s d i v i d e d i n t o t w o mainsec t ions :( 1 ) t h e s t a t i c longi tudina laerodynamic characteris t ics of t h e model and ( 2 ) t h es t a t i c l a t e r a l - d i r e c ti o n a l a e r o d y n a m i c c h a r a c t e r i s t i cs .The s t a l l angleof a t t a c k i s def ined as t he wing angle of a t t a c k a t whicht h e flaw separa tesne ar he wing t i p . T h i s s e p a r a t i o n resul ts i n a sudden lossof t o t a l l i f t and no t i ce ab le po s i t i ve sh i f t in p i tc h i ng moment due t o t h e lossof loading a f t of t he moment c en te r. The maximum l i f t does notalways occur a t

thewing-t ip s t a l l anglebecause hefl ow may re ma inat tached to t h e inboards l a t and f l ap surfaces, therebyproducingaddi t iona l l i f t a t a n g l e sg r e a t e rthan the wing- t ip s t a l l ang le of a t tack. However, the usable range of l i f t isg e n e r a l l y l imited t o t h a t a t thewing-t ip s t a l l ang le because of t headve r see f f e c t s o f t i p flow s e p a r a t i o n o n h e c o n t r o l e f f e c t i v e n e s s of t h e outboard low-speed a i l e r o n or spoiler r o l l - c o n t r o l surfaces. Analysisof t h e wing pressuredata p r e s e n t e d n r e f e r e n c e 5 showed that a l l t h e wing conf igura t ions testedexhib i tedwing- t ip s t a l l behavior . ngenera l , he flaw sepa ra t e snea r t he t i pi n i t i a l l y because of t h e i n a b i l i t y of thehighly hree-dimensionalboundary ayerto remain attached i n t h e presence of t h e a r g e s t a t i c pressure g r a d i e n t s t h a tdevelop a t thehigherangles of at tack. These g r a d i e n t s are due to the combinede f f e c t s of t h e t i p vor t ex roll-up and t h e high oca l ly nducedangles of a t t ack ,which are a func t ion of t h e spanwise load d i s t r i b u t i o no n t h e wing. Based onp o t e n t i a l flaw theory , the spanwise load d i s t r i b u t i o n is a di rec t f u n c t i o n ofthe planformshape,spanwise twist d i s t r i b u t i o n ,a n d local chordwise camberd i s t r i b u t i o n . The s t a l l ang le of t h e cruise wing conf igura t ion tes ted dur ingt h i s i n v e s t i g a t i o n could be increase d by increa s ing t h e wing twist or by droop-ing t h e wing leadingedge increased camber) near t h e wing t i p to reduce t h elocal inducedangles of at tack . However, a t t h e d e s i g n l ig h tc o n d i t i o n s(M = 0 . 8 0 ) , an i n c r e a s e n twist or leading-edgedroop cou ld p o s s i b l y cause t h eformationof local shocks, with a cor re spond ingundes i r ab lereduc t ion n t h edrag-rise Mach number.

The spanwise load d i s t r i b u t i o n of a wing equipped w i t h a h igh- l i f t sys t emis p r i m a r i l y a func t ion of t h e l o c a t i o n ,s i z e , and d e f l e c t i o n of t h e va r ioussystem omponents. For good lowspeed performance characterist ics, it i sg e n e r a l l y desirable to s i z e and de f l ec t t h e variouscomponents so t h a t a l a r g epercentage of t h e t o t a l wing load i s produced by t h e inboardsegment of t h ewing; i n order to reduce t h e s tructural bending moment a t t h e root and reducetheoutboard nducedanglesof at tack. Usual ly, nboard component su rf ac e areasand de f l ec t ions t h a t are p r o p o r t i o n a l l y as g r e a t , i f not g r e a t e r , h a n those of1 2


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theoutboardcomponents are r e q u i r e d . ncon t ra s t , he wing tested d u r i n g h i sinves t iga t ion had p r opor t iona l ly smaller inboardvane and f l ap su r f a ce areas i ncomparison w i t h t h e outboard areas because of the combined e f f ec t s of higha s p e c t r a t i o , low sweep, panw ise ocation of the ra i l ing-edgebreak,and t h edesired l o c a t i o n of the wing wheel well c a v i t i e s . I n s p i t e of the propor t ion-a l l y smaller vanea n d f l a p surface areas inboard, heoutboard nducedanglesof a t t ack could poss iblybereduced by a gr adua l or segmented ncrease n hespanwisedef lect ionof he eading-edge s l a t . Althoughdesirable , s u c h deflec-t i o n s are d i f f i c u l t to obtainmechanical lyand are gene r a l lyno tcons ide r edpr ac t i ca l . The leading-edge s l a t was def lec ted panwise a c o n s t a n t -50, anominaldef lect ionbased on a p r e l i m i n a r y a n a l y s i s of similar cur r en t ly ope r -a t i o n a l h i g h - l i f t s y s t e m s .

Few r e l i a b l e three-dimensiona l ana lyt ica ldes ign methods are c u r r e n t l ya v a i l a b l e t o determine the optimum shape,gap,over lap, and def l ec t i ono f h evarious components of a p a r t i c u l a r h i g h - l i f t s y s t e m h e r e f o r e , e x p e r i m e n t a linves t iga t ion ema ins the on ly re l iab le method. Conversat ionswi th e sea r che r sin nd us tr y who also f l i g h t t e s t f u l l - s c a l e a i r c r a f t i n d i c a t e h a t h e posi-t i o n i n go f s l a t , vane,and f l a p components for optimum performance i s g r e a t l yaffected by Reynolds number. Inadd i t i on ,pe r f o r mance r endsev iden t from wind-tunne l tests a t l o w Reynolds number conditions do notalways emain he same a thighReynolds number f l i g h t t e s t con dit ion s. The adven t of wind tunnels w i t hhigherReynolds number ca pa bi l i t i es , su ch as t h e N a t i o n a lT r a n s o n i cF a c i l i t yunder const ruc t ion a t L a c , w i l l provide a unique oppor tuni ty to perform mored e f i n i t i v e h i g h - l i f t model-scale tests.

L ongi tud ina l Charac ter i s t icsEffe cts of Rey nol ds number .- The ef fec ts of a small change i n Reynoldsnumber on th e untrimmed (h o ri zo n t al t a i l o f f ) longi tudina laerodynamicchar -

acter i s t ics of t h e cruise, climb, p a r t - s p a n l a pa l o n e , u l l - s p a n l a pa l o n e ,f u l l - s p a nf l a pwi th low-speed ai le ro ns , and ful l-sp an f l a p w i t h high-speeda i l e r o ns wing conf ig ur a t ions are p r e s e n t e d n i g u r e s 5, 6 7, 8 , 9,and 10,r e s p e c t i v e l y . The va r i a t ion nReynolds number, based on the reference meangeometr ic chord, was small and anged in va lue from 0.97 t o 1.63 x IO6. T h i ssmall v a r i a t i o n had the expec tedneg l ig ib l e e f f ec t on the aerodynamic charac-t e r i s t i c s of a l l the wing conf ig ura t ion s t e s ted a t anglesof a t t ack below t h es t a l l angle .

The c r u i se and climb ( s l a t def lec ted and f lapsne st ed ) wing conf gura-t ionsdemonst ra ted the typical l i ne a r i n c r ea se in l i f t and p i t ch i ng moment w i t ha n n c r e a s e n h ea n g l e of a t t a c k below t h e s t a l l ang le. The remainin gwingconf igur a t ions w i th both s l a t and f l a p s de f l ec t ed)demons t r a t ed a nonl inea rin c r ea se in l i f t and p i t ch in g moment w i t h a n n c r e a s e n t h e a n g l e of a t t a c kbelow the s t a l l angle .Analys is of the wing pressure data p r e s e n t e d n e f e r -ence 5 showed t h a t this nonl inear behavior was due p t ma r i ly t o t h e f ac t t h a ttheaerodynamic oading on the vane/flapcombinationand rear p o r t i o n of t h emain sect ion emain ed almost c o n s t a n t a s t he ang le of a t t ack i nc r ea sed , whereastheaerodynamic oading on the s l a t and orwardportion of t h e main sec t ioninc r ea sednonl inea r ly .

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Also a s shown in f i g u r e 5 , t h e small increase nRey nol ds number ob tai ne dd u r i n g h i s n v e s t i g a t i o npr oduced nochange in s t a l l angle and a small i n c r e a s ein maximum l i f t a f t e r s t a l l for t h e cruise wing configurat ionsbothwi thnacelles on nd of f . A s shown i n ig u re s 6 to 10, however, a small i n c r e a s ei n Reynolds number pr od uc ed heunexpec tedr e su l to fanove r a l lr educ t ion nboth s t a l l ang le and maximum o b t a i na b l e i f t co e f f i c i en t a f t e r s t a l l f o r h eclimb and for a l l the lapped wing con f igu ra t i ons t es ted . The developmentofasymmetric flow p a t t e r n s b e t w e e n h e r i g h t and l e f t w i ng sresul t ingf romverysmall d i f f e r e n ce s n h e p o s i t i o n s o f h e r i g h t a n d e f t h i g h - l i f t s y s t e m com-ponentscouldhave caused th i s unexpec tedbehavior. The gap s,over laps,andd e f l e c t i o n so f a l l the components were checkedverycareful ly a f t e r t h i s n v e s -t i g a t i o n and were found to be w i t h i ne n g i n e e r i n gspec i f ica t ions . The l e f t wingwas ins t rumentedwi thsur faces ta t ic -pressure taps which could have resu l ted i nasymmetric component deflectionsunderaerodynamic oading;however ,ananalysisof he l a t e r a l data obtained showedno d i s c e r n i b l e p a t t e r n as t o whether her i g h t or l e f t wing s t a l l e d f i r s t . The development of asymmetric or la rgerseparated flow r eg ionsnea r the wing t i p s due to change inReynolds number coulda lso havepossiblycaused this unexpectedbehav ior . The Reynolds number a t t h ewing t ips , based on the t i p cho rd , ranged rom 0.53 to 0.89 x lo6. Even a tt h e s e low Reynoldsnumbers, thee x i s t e n c eo fs i z a b l e e g i o n so f a m i n a r l o w isdou bt fu l because of t h e a r g e pressure g r a d i e n t s t h a t developed on the uppersur f aceo f the leading-edge s l a t and because of heh igh ly u r bu len t flowthrough hes la t /main s lo t e x i t p l a n e .

Very l i t t l e is known a b o u t t h e e f f e c t s o f e i the r dynamic model o s c i l l a t i o n sor highcrossf lowveloci tycomponentson he mechanisms of urb ule ntboundary-l a y e rs e p a r a t i o n a t t h e s e low Reynoldsnumbers. The model had a r a t h e r a r g ef l ex ib le wing and was mountedon a hig hly can t i le vere d model suppor tsystemwhich r e su l t ed n a r ge model andwing- t ipdynamicosc i l la t ionsnear and a f t e rt h e s t a l l ang leo f a t t ack . Dur ing th is nve s t ig a t io n,n o l o w - v i s u a l i z a t i o ns t u d i e s were made t o de te r mine heexac t eg ions of separated f low; herefore ,no def ini t ivee x p l a n a t i o nc a n be given for t heunexpec ted r end n s t a l l ang leand maximum ob ta in ab le l i f t co e f f i c i en t f o r t h e climb and flapped wingc o n f i g u r a t i o n s .

The e f f ec t s o f nac e l l e s on the o ng i tud i na l a e r odyn amic cha r ac t e r i s t i c sof t h e cruise wing conf igura t ion are p r e s e n t e d n f igur e 5 (c ). A t thehigherang le s of a t t a c k , t h e s e da ta show thata d d i n gn a c e l l e s resul ted i n a s l i g h ti n c r e a s en CL and CD and in a v e r y l i g h t o s i t i v e h i f t n C+ Addingnace l l e scanof t encausean nc r ea se n CL, because thenace l les hemselvesproduce a l i f t in cre men t and i n c r e a s e h e loca l wing loads due t o an nc r ea sei n t h e l o c a l inducedangles of attack. A t thehigherangles of a t tack , t h er e l a t i v e l y a r g e increase i n CD was p o s s i b l y caused by the ormati on ofseparated f low egionson henacel les.Such ormationscouldposs ibly bee l imin a ted by properpos i t i on ing of v o r t e xg e n e r a t o r s . A t t h e lower a n g l e s ofa t t a c k , t h e n c r e a s e n CD was much smaller a s a r e su l t of t h e added t h r u s t -i n g force produced by the oe-in of t henace l l epy lon .

Untrimmed ch ar a ct e ri s ti cs .- Summary p l o t s showingcomparisonsof t h euntrimmed lon gi tu din a l ae ro dyn amic cha rac te r i s t ics of a l l th e wing configu ra-t i o n s tes ted are p r e s e n t e d n i g u r e 11 for Rc = 1.63 x lo6. The leading-edges l a t was d e f l e c t e d a c o n s t a n t -50 f o r h e climb conf igu r a t ion and a l l t h e1 4


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f l apped wing co nf ig ur a t ions .T h i sc o n s t a n td e f l e c t i o np r o d u c e de s s e n t i a l l y aconstantspanwise eading-edge camber di f f e r encebe tween hec r u i seandc l imbwing conf igura t ions . As shown i n ig u re 11 ( a ) , for t h e climb c o n f i g u r a t i o n ,t h i s c o n s t a n t camber d i f f e r e n c er e s u l t e d na p p r o x i m a t e l y equal slopes fo r t h eangle-of-attack and C, v e r s u s CL c u r v e s o r h e r u i s e n d climb wingon-f i g u r a t i o n s . An appr oximate 2.5O p os i t i v e sh i f t also occurred i n h ea n g l eo fa t t ack o fz e r o i f t . A t a given CL through a range f .5 t o 1.2, CD was0.02 higher and Cm was i d e n t i c a l o r h e climb, i n compar isonwith he cruisewing conf igura t ions . For t h i s same range f CL, an a n a l y s i s of CD ve r susC L ~ l o t s showed thatb o t hcon f igu ra t io ns had approximate ly he same va lueof0.70 for t he w ing e f f i c i enc y ac t o r e. (D = c D , ~i where CD,o is t h ed r a gc o e f f i c i e n t a t z e r o i f t . The cruise wing c on fi gu ra ti on had an ntrimmedmaximum L/D of16.52 a t CL = 0.68; t h e climb wing co nf ig ur at io n had anuntrimmed maximum L/D of 11.72 a t CL = 0.90.

CL2

) TAeA s shown in f i gu r e 1 1 (a ), CD for he ul l - spa n, ake- of f lap wing on-

f i g u r a t i o n a t a given CL through a rangeof1.2 to 2.2 was approximately 0.01less than CD forhe a r t - span,ake-off lap wing con figu rat i on. A s l i g h t l ylower CD was expected for t h e u l l - s p a n f l a p conf igura t ion ncompar isonwi tht h a t f o r h e p a r t - s p a n f l a p c o n f i g u r a t i o n s as a r e s u l t of thesmootherand moren e a r l y e l l i p t i c spanwise load d i s t r i b u t i o n .A n a l y s i s f CD ve r sus CL2 p lo t sthrough he CL range f .2 to 2.2howed thatbo th ake -of f onf igur a t ionshad approximatelyhe same e va lue of 0.85. Th is ep res en ts a 0.15-increasei n h e e va lue compared t o t h e cruise and climb wing con f igu r a t io ns . ngene r a l , a h i g h - l i f ts y s t e m w i l l improve the flow q u a l i t y , n h a t f l o w sepa-r a t i o n is reduced and add i t iona l e ad ing- edgesuc t ion is recovered so t h a t t h ewing e f f i c i ency is h i g h e r n h e h i g h - l i f t case than ha tof hec lean-wing case.(Also, see ref .13.)Thepar t-span, ake-off lap wing co nfi gu rat io ns had anuntrimmed maximum L/D of 9.61 a t CL = 1.49, ndhe ull-span,ake-off f l a pconf igura t ion had an untrimmed maximum L/D of 10.20 a t a s l i g h t l yh i g h e r CLof 1.55.

A s also shown i n f i g u re 1 1 (a ), h e part- and f ul l -sp an, landi ng f l a p wingc o n f i g u r a t i o n s a t a given CL had a lmo s t den t i ca l CD valueshrough a CLrangeof 1 .0 to 2.6. As f o r hecor r e sponding ake - of fconf igur a t ions , a reduc-t i o n n CD was expected for t h e u l l - s p a n , a n d i n g f l a p wing conf igura t ionscompared with hepar t-span wing config urat io ns.Without d e t a i l e d flowvisu-a l i z a t i o n and more d e t a i l e d s p a n w i s e p r e s s u r e d i s t r i b u t i o n data , n o d e f i n i t i v eexplana t ioncan be given for t h e lack of r e d u c t i o n n CD. Perhaps the f u l l -span, l and ing f l a p wing configurat ion had a r eg ion of separated flowwhichtended t o i nc r ea se hedr agof he model. However, such a r e g i o n i s n o t e a d i l yapparentfrom an a n a l y s i s of t h e l imi ted wing pressure d i s t r i b u t i o n da ta takend u r i n gh i snves t iga t ion . An ana lys i s of CD ve r sus CL2 p lo t s showed thatbo th and ingconf igura t ions had an e valueof ppr oxi mat ley 0.89, whichr e p r e s e n t s a f u r t h e r n c r e a s eo f 0.04 compared wi th heva lue o r he ake - of fconf igur a t ions .Both and ingconf igura t ions had an approximateuntrimmed m a x i -m u m L/D of 7.55 a t an approximate CL of 2.0.he part-span,anding ndful l-span, take-off f lap-wingonf igura t ions had ne ar l yd e n t i c a le r s u sCL c u r v e s o r a CL rangeof 0.7 t o 2.5,which implies t ha tbo thconf ig -

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u r a t i o n s n c u r h e same trim d r a gp e n a l t i e s . ( T r i m drag is de f ined a s t hedr agincrementdue to t h e h o r i z o n t a l -t a i l l i f t r eq ui re d to trim t h e a i r c r a f t . )

Comparisons of th e untrimmed lo ng it ud in al a e r o d y n a m i cc h a r a c t e r i s t i c s ofthe our ake-of f and lan din gf la p wing conf igu ra t ions are p r e s e n t e d nf i g -ures 11 (b) and 11 ( c ), r e spec t ive ly . For t h e take-off f l a pc o n f i g u r a t i o n s a t agiven CL through a range of 0.2 t o 2.1, the ul l -spa n f l a p and full-span f l a pwith low-speed a i l e r o ns wing conf igur a t ions had s l i g h t ly lower CD and, here-fo re , higher L/D values han he or respondingva lues for t h ep a r t - s p a n l a pand ul l - span lapwi thhigh-speedaileron s wing con figu rat i ons . The ful l-sp anf l apwi thh igh- speeda i le ron s wing conf igu ra t ion pr oduceds l igh t lyhigher neg-a t i v e C than he ul l - span f l a p with low-speed ai le ro ns wing config urat io n.T h i s r e s u l t was expected because of t he educed oad ingnea r he ipsof hewingequippedwithoutboard low-speed a i l e r o n s . For t he and ing l apconf igu-r a t i o n s t e s ted a t a given CL through a range of 0.7 t o 2.6, a l l theconf igura-t i o n s t e s t e d had approximatelyhe same CD and L/D values . The t r en ds nt h e C c h a r a c t e r i s t i c s for t he and ing f l a p wing configurat ions were similart o thoseobserved or he take-off f lap wing configurat ions. For bot h he take-of fand and ing lap wing conf ig ura t ion s , he hree ul l - sp an f l a p wing configu-ra t ionsproduced more nega t ive C, than t h e part-span f l a p wing configurat ion;t h i s r e s u l t implies higher trim d r a g p e n a l t i e s f o r h e f u l l - s p a n f l a p wingconfigurat ions.Thesehigher trim d r a gp e n a l t i e sc o u l de a s i l yo f f s e ta n yimprovements i n L/D obta ined s ing ul l - span f l aps .

For t h i s model, t h e o v e r a l l trim d r a gp e n a l t i e sm i g h t be reduced by movingt h e moment re fe re nc ec e n t e r u s u a l l y h e same as t h e a i r c r a f t center-of-gravityl o c a t i o n ) u r t h e r a f t to r educe henega t ive C . However,moving th e en te r-o f - g r av i ty oca t ion a f t without a co rr es po nd in g af t movement of therear-wheelloca t ioncanadve r se lyof f se t henose - whee l t e e r ing o r ce s . ngene r a l , amore d e s i r a b l e way of r educ ing heove r a l lnega t ive Cm value is to d e s i g n h ewingand f l a p system so t h a t t h e maximum l i f t w i l l be genera ted by the nboardsegment of t h e wing. Th is method w i l l reduce henega t ive Cm value by movingth e wing cen te r -of -p ressure oca t ion thepoint hroug h which th e wing re su l t an tforce ac t s ) forward toward t h e model moment referencecente r .Thisapproachw i l l requi re a rge-percent -chord,highlydef lec ted f l a p s inboard t o p ro du ce l i f tandhighly deflected leading-edgedevicesoutboard to i n c r e a s e h e s t a l l angleand t h e maximum ob ta in ab le l i f t . As previouslydiscussed,higher nboard load-i n g w i l l a l s o have he added advan tageofr educeds t r uc tu r a l root bendingmoments and of lower inducedangles of a t t ack outboard.

Trimmed character i s t i c s .- T h e e f f e c t of h o r i z o n t a l - t a i l d e f l e c t i o n o n t h el o n g i t u d i n a la e r o d y n a m i cc h a r a c t e r i s t i c s of t he10 wing conf igur a t ion s nves t i -ga ted are p r e s e n t e d n i g u r e s 12 to 17. The lo ng i tu din a l trim c h a r a c t e r i s t i c s(Cm = 0 ) , determined by i n te r pol a t in g he e x p e r i m e n t a l da ta curves t o o b t a i ndata a t incrementa l t a i l inc idences , a re a l so presented oreachconf igura t ion.The maximum trim CL is de f ined as t h eh i g h e s tv a l u e of CL obta ined prior ton e u t r a ls t a b i l i t yo f h e model (aC&CL = 0 ) . The angleo f a t t a c k f o rn e u t r a ls t a b i l i t y o c c u r r e d s e v e r a l d e g r e e s p r io r t o thewing- t ip s t a l l angleof a t t a c kf o reachconf igur a t ion t es ted .

The tail-offper formance da ta p r e s e n t e d n i g u r e s1 2 t o 17 show tha t hec r u i s e and anding f l a p wing configurations had a l a r g e r p o s i t i v e s h i f t n Cm16


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a f t e r t h e s t a l l angle han d i d t h e climb and ake-offf lap wing conf i gurat ions .T h i sp r obab ly nd ica te s ha t he cruise and l an di ng f l ap wing con f igur atio ns hadla r ge r r eg ions o f separated f lownea r he wing t i p s and,consequently, had moreforwardcenter-of-pressure ocations han hose or hecorresponding climb andtake- of f f lap wing conf igura t ions .The tail-onperformance data showed th at th e cr ui se and la ndi ng f la p wingconf i gura t ions had a s l igh tun f avor ab lepos i t ive h i f t nose - up ) n C, andt h a t t h e climb and ake-offflap wing con fig ura ti ons had a s l i g h t f a v o r a b l ene ga t i ve h i f t (nose-down) i n C, a f t e r h e n g l e o rn e u t r a l t a b i l i t y . ngenera l , he two best approaches to reduceandpossibly to e l i m i n a t e h i sunfavorablepitch-up are (1) t o inc r ea se the h o r i z o n t a l - t a i l e f fec t iveness and( 2 ) to reduce t h e separated f low egionnea r he wing t i p s . Greater hor izonta l -t a i l e f f e c t i v e n e s s c a n be achieved by inc rea sin g he sur fac e area of he cu r r en tlow-tail arrangement,which may i ncu radd i t iona l cruise dr agpena l t i e s , or byincreas ing he t a i l moment-arm length w i t h t h e use of a high T-tai l arrangement.The high T - t a i l would a l so be i n a regionofhigher local f low han ha t for t h ec u r r e n t low-tail arrangement,which would require a smaller t a i l su r f ace areato trim the model and, herefore , would incur less trim dragpena l ty .Neither ofthese two approaches w i l l e l imina te heunfavorable ipf low-separa t ion reg ionwhich adv ers e ly a f f ec t s t h e r o l l - c o n t r o l e f f e c t i v e n e s s of the low-speed a i le ron s .Perhaps a bet te r approach is to reduce andpossibly to e l imina te the reg ionofseparated flow near t h e t i p s by inc r ea s ing he outboard s l a t a nd n bo ar d l ape f f ec t ivenes s . As pevious ly discussed, th i sapproach could resul t i n h e morefavorablecond i t ionof flow separa t ionne ar he wing-body ju nc tu re . However,i n i t i a l n b o a r d f l o w s e p a r a t i o n would a d v e r s e l y a f f e c t h e h o r i z o n t a l - t a i le f f ec t ivenes s o f a high T - t a i l arrangement a t theh igherangles of a t t ack ,espec ia l lynear s t a l l .Summary comparisonsof he trim performance for he wing con f ig ura tio nstes ted are presented n igure18 . A t abulated summary of the trim drag pen-a l t i e s and the maximum trimmed CL and L/D performancevalu es or ach wingconf igura t ion tes ted is pr e sen ted n t ab le 11. The maximum trimmed performancedata f o r h e cruise , climb, andpart-span, t a k e - o f f and l an di ng la p wing on-f igurations compare favorably w i t h theperformance data obta ineddur ing a pre-v ious nves t iga t ion as reported i n e f e r e n c e 3 . The trimmed maximum L/Dva lues f o r he cruise and climb wing conf igurations were o n l y s l i g h t l y lowerthan t he i r correspondinguntrimmedvalues;however, t h e valueof CL a t which

the maximum L/D occurred was approximately 0.1 4 higher . As previous ly d i s -cussed, the u l l - span , ake-of f lap wing conf igura t ion had s l i g h t l y bet t e runtrimmederformance (lower CD and igher L/D a t a given CL) than t h epart-span, ake-offwing onfi gura tion . However, as shown i n i gu re18(a) andgiven n t ab le 11, t h e trimmed data show tha t thepar t - span , ake-of fflap wingconf igura t ion had s l i g h t l y bet ter trimmed performance han he ull-span, t a k e -o f f f l a p c o n f i g u r a t i o n s , a n d t ha t thepar t-span, anding f l a p conf iguration hadconsiderably bet ter trimmed performance han he ull-span, anding lap con-f iguration. The loss in heperformance of t h e f u l l - s p a n l a pc o n f i g u r a t i o n swas due to the h igh trim p e n a l t i e s i n c u r r e d i n t h e form of hor izonta l - ta i l d ragand downloads. The trim dragpena l ty for the u l l - span , take-off f l apconf ig -u r a t i o n was twice as g r e a t as t h a t for t h e part-span, ake-off f l a p config-uration and was nea r ly 3 times as g r e a t f o r t h e cor responding andingf lapconf igura t ions .17


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Comparisons of the trimmed longi tud ina laerodynamic characteristics of t h efour take-off and t he four l an d in g f l a p wing con f igu r a t ions are p r e s e n t e d nf i g u r e s 18 (b) and 18 (c), r e spec t ive ly . The differencesbetween the trimmed takeoff f l ap pe r f o r mance cha r ac te r i s t i c s of t h e four f lapped wing conf igurationstested were very small. The average trimmed maximum L/D was 9.2 a t an verageCL of 1.53. However, t h e trimmed landing f l a p performance characterist ics oft h e part-span f l a p and ull-span f l a p w i t h low-speed ai l er on s wing configura-t i o n s were almost i d e n t i c a l and were cons iderab ly better than th e performanceof t h e fu l l - span f l a p andfull-span f l a p wi t h high-speedaile rons wing config-ura t ions . T h i s resul t was expected because of an ncrease nnega t ive C, causeby the a f t movement of the wing cent er-of -pres sure ocat ion as more of t h e outbopor t ion of t h e wing was loaded.

The full-span f l a p wi th high-speed a i le ro ns wing conf i gura t ion wi th bothtake-off and anding f l a p s e t t i n g s had s l i g h t l y lower maximum CL va lues hant h e correspondingpart-span f l a p wing conf i gura t ions . nadd i t ion , t h e f u l l -span f l a p w i t h low-speed a i l e r on s wing con f igu r a t ion wi th both take-off andlanding f l a p s e t t i n g s had s l i g h t l y higher maximum CL va lues han t he corres-pondingpart-span f l a p wing conf igurations . The full-span, take-off f l a p wingconf igura t ion had a higher maximum CL va lue han he other take-off wing on-figurations; however, t h e full-span, anding f l a p wing conf iguration had a lowermaximum CL value han th e par t-span f l a p conf igurationandonly a s l i g h t l yhighervalue t h a n t he fu l l - span f l a p w i t h high-speedaile rons wing configura-t ion .Comparativeana lys is of the trimmed performance character is t ics of t h efour take-off and four landing f l a p wing conf igurationssugges t s t h a t t h e f u l l -span f l a p with low-speed a i l e r on s wing con f igu r a t ion had s l ig h t l y bet ter trim-med performance characteristics than the other th r ee con f igu ra t ions . n addi-t i o n , t h e trim performance of the fu l l - s pan f la p wing conf i gura t ion was s l i g h t l yworse thaneven t h a t of the pa r t - span f lapconf igura t ion because of high trimdrag p e n a l t i e s . These tes t resul ts f u r t h e r r e i n f o r c e t h e g e n er a l lyacceptedphilosophy for t h e des ign of h i g h - l i f t systems: t h e more t o t a l l i f t generatedby the inboardpor t ion of the wing, the bet ter the ove r a l lper formance of t h es y s em.

E f f e c t of landing gear.- The effects of the landinggear on t h e longi tu-dinalaerodynamic character i s t ics of the ake-off and landing f l a p wingcon-f i g u r a t i o n s tested are pr e sen ted n igu r e 19. These data show t h a t a t a givenC L , adding t he landinggear had n e g l i g i b l e e f f ec t on C, bu t resulted i na naverage ncrease n CD of 0.01 4 for t h e take-off f lap wing conf igurationsanda s l i g h t l y smaller average ncrease of 0.010 for the l and ing f l a p wingconfigu-r a t i o n s . A t a givenangle of a t t ack , adding the landinggear resul ted i na napproximate CL loss of 0 .0 3 for t h e take-off f l a p wing conf igurations and aloss of 0.06 for the landing f l a p wing conf igurations . These l i f t losses wereexpected because of t h e i n t e r a c t i o n of gear-generated ow-energy w a k e s w i t h t h ehigh-energy flow through t h e main/vaneandvane/flap slots. Ingenera l , t h i si n t e r a c t i o n reduces the energy of t h e flow through t h e s lots and ther eby educ est h e l i f t incrementgenerated by the inboard f l a p system.

Effec t of leading-edgesla ts . - The effec ts of t h e l e ad ing - edge s l a t deflec-t ion on the longitudinalaerodynamic characterist ics of the part- andfull-spanf l a p wing conf igurations with both take-off and anding f l a p s e t t i n g s are pre-sen ted n igu r e 20. These data show the tremendous effects of d e f l e c t i h g t h e18


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leading-edge s l a t on t h e maxjmum s t a l l angleandon he maximum CL c a p a b i l i t yof the ake-of f and lan din g lap wing conf i gura t ions .Nes t ing he s l a t r e s u l t e dinanapproximate18.5- perce nt educt ion n he maximum ob ta in ab le CL and napproximate educt ion n CD of 0.010 a t a given CL. The wing pressure datap r e s e n t e d n r e f e r e n c e 5 show that he nboard eadi ng-ed ge s l a t segment(between uselageside-of-bodyand railing-edge break s t a t i o n s ) carr ied lessload han he outboard s l a t segment a t a givenangle of a t t ack . This was duei n p a r t to t h e f a c t t h a t t h e thickness and camber d i s t r i b u t i o n s o f h e n b o a r ds l a t were gr ea te r han hose of theoutboard s l a t and i n p a r t to t h e f a c t t h a tboth nboardandoutboard s la ts were d e f l e c t e d h e same amount. These combinedf a c t o r s r e s u l t e d nh i g h e r e f f e c t i v e e a d i n g - e d g e camber inboard hanoutboard.However, for th i sh igh- aspec t - ra t io wing conf igura t ion , less leading-edge camberi s needed nboardbecause of t h e small percentchord of the nboardf lapsystem.The e f f ec t ive camber can be reducedandposs ib lyhigheroverall maximum CLva luescan be obtai ned by decrea sing he nboar d s l a t d e f l e c t i o n or by replac-ing he s l a t with a properlycontouredvariable-camberKrueger (VCK) leading-edgedevicesuch as tha t used by Douglasandrepor ted nreference 4.

E f f e c t ofgroundheight. - The e ff ec t s of groundhe ighton he ongi tud ina laerodynamic character i s t i c s of the part- and fu l l - span f lap wing conf igura t ion swithboth take-off and landing f l a p s e t t i n g s are pre se nt ed n ig ur e 21. Theh o r i z o n t a l ta i l s were onduring hesegroundproximity tests . The data show theexpected i n c r e a s e n CL, r e d u c t i o n n CD and pos i t ive s h i f t i n Cm with adecrease i n groundheight.Thesechanges nperformanceoccurbecause t h ecushioning effec t of hegroundsuppresses heformation of the wing vortex,wi th a r e s u l t i n g decrease i n wing downwash and i nc re as e n if t- cu rv e slope.Both par t- and fu ll -s pa n f l ap wing con f ig ura ti ons had anapproximate ncreasei n CL of 0.08.The take-off f l a p wing co nf ig ur at io ns had an pproximater educ t ion n CD of 0.02nd a n e g a t i v e h i f t n C m of 0.06. However, thelanding f l a p wing con fi gu ra ti on s had a g r e a t e r e d u c t i o n n CD of 0.04 andn e g a t i v e h i f t n Cm o f 0.08. The n e g a t i v e h i f t n Cm will probablyr equ i r eg r ea te r ho r izon ta l - t a i l nc idenceang le s to trim t h e model, which w i l l ,i n u r n , n c r e a s e h eo v e r a l ld r a g due to t h e i n c r e a s e n trim drag.

The effects of groundheight on the ongi tudi nalaerodynamic charac te r -i s t ics of t h e par t-span, anding f l a p wing configurationwith heground andf l i g h t s p o i l e r sd e f l e c t e d 4 5 O and 60 are presented n igure21(c) .Thesedata show an p proxi mate ncrea se n CL of 0.11, an i n c r e a s e n CD of 0 . 0 2 ,and a p o s i t i v e h i f t n C, of 0.06. The in cr ea se n CD and thep o s i t i v es h i f t n Cm were unexpectedandopposite to the rendsobserved for t h e sameconf igura t ionwi th he spoilers undeflected.These data also show thatd e f l e c t -ing heg ro u nd and f l i g h t spoilers 45O re su lt ed n an approximatenet loss i nCL of1 .2 ; naddi t ion , ncreas ing the d e f l e c t i o n to 60 r e s u l t e d na n addi-t i o n a l loss of 0.3 i n CL.

L a t e r a l - D i r e c t i o n a l C h a r a c t e r i s t i c sE f f e c t of s ides l i p anq1e.-The ef fec ts of s ides l ip angl e on th e la tera l-d i r e c t i o n a l a e r o d y n a m i c c h a r a c t e r i s t i c s o f a l l the wing co nf igu ra t i ons nves t i -ga ted , except he fu l l - span f l a p withhigh-speed a i le ron wing conf i gura t ion wi tht a k e - o f f l a ps e t t i n g s , are pr e sen ted n igu r e s 22 to 27. The s t a t i c l a t e r a l -

19


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d i r e c t i o n a l s t a b i l i t y d e r i v a t i v e s are presented for each conf igu r a t ion testedfor angles of attack below s t a l l and were computed from t he l a t e r a l body-axisdata obtained a t s ides l ip angles from -loo to 5O. The h o r i z o n t a la n dv e r t i c a ltails were onduring a l l runs to determine t h e e f f e c t of s ides l ip . The c r u i s ewing conf iguration had a geometr ic d i he d r a l angle of 5O; t h e h o r i z o n t a l t a i l shad a geometr ic d i he d r a l angle of loo. These dihedra l angles resulted i n s tablel a t e r a l - d i r e c t i o n a l characterist ics for each conf igu r a t ion tested.

A s shown i n f igu r es 22(b) and23(b), t h e climb wing configuration had lessl a t e r a l - d i r e c t i o n a ls t a b i l i t y h a n the cruise wing conf iguration. The e f f e c t i v edihedra l parameter C z g showed an pproximate 0.001 p o s i t i v e s h i f t . T h is reduct i o n i n l a t e r a l s t a b i l i t y is be l ieved to have resulted pr imar i ly from a decreasei n geometric d ihedra l which was caused by def lec t ing t h e leading-edge s la t sspanwise a t a cons tan t ang le of -5OO.

The e f f e c t i v e d i h e d r a l parameter Czg decreased r ap id ly from approximately-0.002 a t -4O to approximately -0.006 a t loo angle of a t t a c k for a l l of t h etake-off and landi ng lap wing conf i gura t ions tested, i n d i c a t i n g an inc r ea se nl a t e r a l s t a b i l i t y with increasedangle of at tack. A t a givenangle of a t t ack ,the va lue of C for each landing f l a p wing conf iguration was o n l y l i g h t l ymre negative han t h e value for the corresponding take-off f l a p wing configu-r a t ion . A t a given CL, the re fore ,he a lue of C for each take-off f l a pwing configuration was much more nega t ive han th e va lue for the correspondinglanding f l a p wing conf iguration. The landing f l a p wing configuration wasexpected to h a v e s l i g h t l y less l a t e r a l - d i r e c t i o n a l s t a b i l i t y t h a n t h e "corre-sponding take-off f l a p wing configurationbecause of t h e inboard s h i f t i n wingloading as t h e f l a p d e f l e c t i o n was increased. The d i r e c t i o n a l - s t a b i l i t yparameter CnB, which is governedprimarily by t he v e r t i c a l t a i l area andloca t ion , anged nva lue from 0.002 to 0.004 for each wing conf igurationtested. The side-f orce parameter Cy0, which i s governedprim aril y by th ef use lage shape, had an averagevalue of -0.02 for each wing conf igurationtested. These values of both t h e d i r e c t i o n a l - s t a b i l i t y a n d side-force param-.eters are typical for this type of wide-body transport.

10

26

Effect of law-speedailerondef lection.- The ef fec ts of low-speed a i l e r o nde f lec t ions on the la te ra l -d i rec t i ona l ae rodynamic characterist ics of t h e part-span f l a p wing configuration w i t h both take-off and landing f l a p s e t t i n g s arepresented n f igu re 28. These data show tha t nega t ive e f t -a i le ron-only deflec-t i o n s produced approximately twice t h e r o l l i n g moment a s p o s i t i v e d e f l e c t i o n sfor the take-off f l a p wing configurationandnear ly 3 times as much for thelanding f l a p wing conf iguration. The highernegative left-aileron-only-deflec-t i o n s produced mre nega t ive Ci for the landi ng lap wing conf igura t ion hanfor t h e take-off f l a p wing configuration.Unexpectedly,however , the h ighe rpos i t ive e f t - a i l e r on - on lyde f lec t ionsp r oduced l igh t ly more p o s i t i v e Cz fort he take-off f l ap wing confi gurat ions han for t he Landing f l a p wing configu-r a t i o n .D i f f e r e n t i a ld e f l e c t i o n s of the r i g h t and l e f t a i l e r o n s produced t h e20


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e x p e c t e d a d d i t i v e r e s u l t s for both ake-of fand andingfla p wing configura-ti on s. None of he data presented for theva r ious f l a p and ow-speed a i l e r o nwing combinations tes ted show s i g n i f i c a n t effects on he ongi tud ina landl a t e r a l aerodynamic coeff ic ientso t h e r h a n C 2 .E f f e c t of s p o i l e r d e f l e c t io n . - The e f f e c t s o f t h e l e f t o u t b o a r d r o l l - c o n t r o lspoiler d e f l e c t i o n on the ong i t ud ina l and l a t e r a l ae r odynamic cha r ac te r i s t i c s ofthe f u l l - span f l a p wing conf iguration w i t h both take-off and andingf lap set-t i n g s are p r e s e n t e d n i g u r e s2 9 ( a ) t o 2 9 ( f ) . The combined effec ts of the l e f tou tbo a r d f l ig h t and r o l l - con t r o l spoiler def lections on heaerodynamiccharac-t e r i s t i c s of thefu l l - span anding f l a p wing conf iguration are p r e s e n t e d n f i g -ures 29 (9) to 29 i) A summary figure showing thechange n t h e aerodynamicc o e f f i c i e n t s Cm, CL, and C 2 withncreased spoiler d e f l e c t i o n is a lso pre-sen ted oreachconf igura t ion tes ted. These summary figures show t h a t a r g ed e f l e c t i o n s of the l e f t outboard o l l -contro l spoiler f o r h ef u l l - s p a n f l a pconf igurationproducedchanges n C l a s g r e a t as thoseproduced by di ff er en -t i a l low-speed a i le ron def lec t ions for hecor responding par t - span f l a p configu-ra ti on . However, th e ar g e spoi ler d e f l e c t i o n s also unloaded he outboard seg-ment of the wing,which resulted i na n expected loss of CL and a p o s i t i v es h i f t n h. D u r i n g c t u a l l i g h t p p l i c a t i o n , h i s loss i n CL and s h i f t i n

in groundproximity.couldadverse lyaf fec t hehandl ing q u a l i t i e s of t h e a i r c r a f t , e s p e c i a l l y

The negative C l produced by the o l l - contr o l spoiler def lec t ions bovel o o was g r e a t e r for t h e take-off than or t h e landing f l a p wing configuration.A s shown i n f igu r e s 29 ( f ) and 29 i) henega t ive C 2 produced by the combinedf l i g h t and r o l l - c o n t r o l spoiler d e f l e c t i o n s was approximately 3 times g r e a t e rthan henegative C 2 produced by roll-control spoiler de f lec t ionson ly .Accordingly, he hange n CL was also approximately 3 times grea te r ,and t h echange in Cm was approximately twice as gr ea t . The d a t a pr e sen ted n he sef i g u r e s a l so show an almost l inea r nc r ea se nne ga ti ve Cn wi thangle ofa t t a c k f o r t h e va r ious spoi ler d e f l e c t i o n s tested. T h i s resul t was expecteddue to the inc r ea se nd r agof he l e f t wing caused by the separated f lowgenerated by t h e def lected l e f t spoi ler .

SUMMARY OF RESULTSAn i n v e s t i g a t i o n was conduc ted n heLangley 4- by 7-Meter Tunnel todetermine he s t a t i c long i tud ina land a te ra l -d i rec t iona laerodynamic charac-t e r i s t i cs of anadvancedaspect- ra t io-10 supercr i t ical-wing ranspor t modelequippedwith a full-span eading-edge s l a t and part-spanandfull-spandouble-s lo t ted t r a i l ing-edge f l aps . Th is wide-body tr an sp or t model was a lso equippedwith spoi ler and ai l eron oll- contr olsur faces , low- throughnace l les , andinggear, andmovable hor izontal tails. The f o l l o w i n gs i x basic wing conf igura t ionswere tes ted dur ing h i s nves t iga t ion : 1 ) r u i s e ( s l a t s and f l aps n e s t e d ) ,(2) climb (s l a t s deflected and f l aps n e s t e d ) , ( 3 ) par t - span lap , ( 4 ) full-spanf l a p , (5 ) fu l l - span f l a p wit h low-speed ai le ro ns ,and ( 6 ) full-spanwithhigh-speed a i le rons . Each of the ou r flapped wing conf igurations was tested withthe leading-edge s l a t s and th e ra i l i ng-ed ge f l aps deflected t o s e t t i n g s repre-s e n t a t i v e of both take-off and andingcondi t ions . The re su l ts of t h i s inves t i -ga t ion are sunnnarized as fo l lows:

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1. The small v a r i a t i o n i n t e s t Reynolds number from 0 .97 to 1 .63 x 106 hadn e g l i g i b l e e f f ec t on t h e aerodynamic characteristics of a l l t h e wing configu-r a t i o n s tested a t ang le s of attack below the s t a l l angle .2. A l l t h e wing conf igurations tes ted exhib i tedwing- t ip - s ta l l behavior

followed by a r e d u c t i o n n o n g i t u d i n a l s t a b i l i t y .3. For t h e take-off f l a p c o n f i g u r a t i o n s a t a given untrimmed l i f t coeffi-c i e n t CL through a range of 0.2 t o 2 .1 , the u l l - span f l a p and ull-span f l a pwi t h low-speed a i l e r o ns wing con f i gu r a t ions had s l i g h t l y lower d r a g c o e f f i c i e n tCD and, therefore, higherva lues of l i f t - d r a g r a t io L/D va lues han t h ecorrespondingvalues for the par t-span f l a p and full-span f l a p w i t h high-speeda i le ron s wing conf igu ra t io ns . For the landing f l a p conf igu r a t ions a t a given

untrimmed CL through a range of 0.7 to 2.6, t h e four lapp ed wing configura-t i o n s had approximately the same CD and L/D va lues . For both take-off andland ing l apconf igu r a t ions a t a given untrimmed CL, t h e three f u l l - s p a n l a pwing configurationsproduced more negativepitching-momentcoefficient C, thanthe par t-span f l a p wing configurationand, therefore, incur redhigher trim dragp e n a l t i e s .

4 . Because of trim d r a g p e n a l t i e s , t h e differencebetween t h e trimmed per-formance characteris t ics of the four flapped wing conf igurations w i t h take-offf l a p s e t t i n g was small and t he average trimmed maximum L/D was 9.2 a t anaverage CL of 1.53. With landing f l a p s e t t i n g s , the trimmed performances oft h e part-span f l a p and full-span f l a p wi t h low-speed ai le ro ns wing conf igura-t i o n s were almost i d e n t i c a l and were cons iderab ly better than t he performancesof t h e fu l l - span f l a p and full-span f l a p wi th high-speed ai le ro ns wing configu-r a t i o n s . A compara t iveana lys is of the trimmed performance of the four flappedwing conf igurationssugges t s t h a t t h e fu l l - span f l a p w i t h low-speed a i l e r o n swing conf igur a t ion had s l i gh t ly bet ter trimmed performance han t h e other threef l a p conf igura t ions .

5. Adding the landinggear a t a givenangle of a t t a c k resulted i n a napproximate CL loss of 0.03 for t h e flapped wing conf igurationswi th take-offf l a ps e t t i n g s a n d a CL loss of 0.06, w i t h l a n d i n g l a p e t t i n g s .6 . For t h e part- and full-span f l a p wing conf igurations w i t h ei ther take-off or landing f l a p s e t t i n g s , n e s t i n g the leading-edge s l a t r e s u l t e d na napproximate18.5-percent educti on n t h e maximum obtainable CL and napproximat e 0'.010 re du ct io n n C a t a given CL.7. Ground-proximity tests of t h e pa r t - and fu l l - span f l a p wing configu-r a t ionsw i t h ei ther take-off or landing f l a p sett ings showed t h e expected rendof a n n c r e a s e n CL, r e d u c t i o n n cD# and pos i t ive s h i f t i n Cm w i t h adecrease i n ground h e i g h t .8. The climb wing co nf ig ur at io n had less l a t e r a l - d i r e c t i o n a l s t a b i l i t ythan the cr ui se wing conf ig urati on. The l a t e r a l - d i r e c t i o n a ls t a b i l i t y of eachf lapped wing conf igu rat ion w i t h landing f l a p s e t t i n g s was s l i g h t l y less than t h es t a b i l i t y of the cor responding conf igura t ion wi t h take-off f l a p s e t t i n g s .

22


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9. For thepar t-span wing conf i gurat ion,nega t ive e f t -a i le ron-onlydef lec -t ionsproduced pproximately twice the oll ing-mom ent oeff ic ient C l aspos i t ivede f lec t ions f o r heconf igu r a t ionwi th ake - o f ff l apse t t ings ,andnear ly 3 times as much w it h and in g f lapse t t ings .10. For t h e f u l l - s p a n f l ap wing conf igurationwitheither ake-off orl a n d i n g f l a p s e t t i n g s , a r g e d e f l e c t i o n s o f h e e f t outboard r o l l - con t r o lspoilerproducedchanges n C l as g r e a t as those produced by d i f f e r e n t i a llow-speed a i le ron def lec t ions for he cor responding par t - span f lap wingcon-f igu r a t ions .La r gede f lec t ionso f he o l l - con t r o l spoi lers a lso r e s u l t e d nan unfavorable loss of CL, a p o s i t i v e s h i f t n Cm, and an i n c r e a s e nnegative Cn.

Langley Research CenterNationalAeronautics andSpaceAdministrationHampton, VA 23665January 30, 1981

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APPENDIX A

LONGI TUDI NAL STABILITY-AXIS AND LATERAL BODY-AXIS DATAThe force and moment data , pr e sen tedg r aph ica l ly nf igu r e s 5 to 29, arepresented n t abular form in h is ppe ndi x . The lon gi t ud i na l data CL, CD,, and L/D (CL, CD, CPM, and L/D, r e s p e c t i v e l y ,n tabular form) are refer-enced to the t ab i l i ty - ax i s y s tem; he l a t e r a l d a t a C 2 , Cn, and Cy (CRM,CYM, and CSF, r e spec t ive ly , n abu la r o r m) are re fe renced to thebody-axissystem.These data were obta ineddur ing Tes t 198 conducted n heLangley 4-by7-MeterTunnel.

24


28/184

L W l G I T U D I N A LS T A B I L I T Y - A X I S A ND L A T E R L L O D Y - A X IS D A TAR C T A - D E GL P H A * D E QL CD CPMRY

.oo -4.04 -e1458 .0270.1407 . O O 2 O. o o -1.95

. O O .OT -2620 -0264 - . l l 4 5 -0019.oo 2.13 -4607 .0299 - . O B 6 5 -0036.oo 4.16- 0 0 6.23 ,6659 .on4 -.0611 . n o 0 5- 8 5 5 4 - 0 4 6 2 -.0317 -0021

-08460247 -.125R - 0 0 3 4

-~~.oo 8;ZS 1.0087 .Os18 - . a 0 1 5 .000710.30 1.17780751 . o m 6 .0019oo.01.02 12.28.1724 -1600 .1442 . 0 0 0 5- . o o 14.29.2084 .PO56 .PO82 - . 0 0 0 816.31.1136 .3040 .1599 .0192- . o o.oo 20.22 1.0822 -3940 .le57 -017418.30 1.1124 ,3547 .17650191.01 22.23 1.0684 - 4382 ,1994 mol54

CY*.0010.0008.0008.0011.0010.0014.0008. O D 1 2. 0 0 0 1.0004-0043. 0029

L O N G I T U D I N A LS T A B I L I T Y - A X I S A ND L A T E R A L D D Y - A X IS D A TAR F T A * D E GL P H A e D E GL CDPMRM CYU

. o o. o o- . oo. o o.oo.oo.Ol.oo-.01-.01.02.01

-4.02 "1293 -0271-1.96.08

e0771 ,02432.15

.2720 .02554.18 -4822 -02946.22

,6714 ,03718.28 1.0412 - 0 6 1 2.e6010478.10.30 1.2056 .075112.28 1.2314 -149314.29 1.2847 .1900

L O N G I T L I D I N A L T A B I L I T Y - A X I SP F T A v D E GL P H P I D E GL

.oo.oo -4.04 "1239-2.00.oo .0778-.OD

-06 28412.15- . o o -4869.oo 4.186.24 n6.930. o o 8.26.0516 A9750-.01 10.30 1.216803 12.28 1.2499.Ol 14.28 1.2582.01.oo 18.21.179316.22.2775.01 20.22 1.1498.02 22.25 1 0987e 0 3 24.19 1.1165L O N G I T U D I N A LB T L B I L I T Y - * X I S

RFTA .DEG ALPHA.DEG CL.oo -1.970656.oo -07.oo -259.2.OO

2.14 .4706.oo 4.186.24 -8637-6664.01.01

8.23 1.036410.29.1899.01-01 12.29 1.242514.30 1.3001

-.14070009"1117 . 0 006-.OR97 ~0025-.0641 .0008-.0046 - e 0 0 0 2-.0355 - 0006

.0364 -0014.1343 e0097.PO15 .0112.E632 -.0022-1516 , 0 0 4 8

-.~zno .0018 ,0009,0005.0002. 0 0 1 0.0007,0010- 0 0 0 4.0012.0051- 0 053-.0008.0011

AND LATERAL SODY-co CPY

.0260 "1365,0232 -.1256.0245 "1077.0288 -.0880-0363 "0615-0467 "0339e0755 -0390.1465 ,1357.2060 ,2044-2711 .ZOO1.3568 -1509,3966 ,1750.4390 e1707.4888 .1958

.0601 - . o o ~ n

, A X I S D A T ACRU CYU.0031 -0003-0024 .0005-0022 ,0003.0029 ,0007.0013 .0010-0009 -0009.OOO3 ,0010. 0003 ,0008.0115 -0070

- .0012 -.0001-.0010 - .0012.0029 - 0 0 0 3. 0043 -.0029-0016 "0014-0013 "0016

P N D L A T E R A LR O D Y - A X I S D A T ACD CPYRY

.0244 "1360 .0020.0284 "1144 .0006-0351 -.0831 -0018

.0611 -.Ole6 .0005-0472 -.OS39,0009

.0791 -0185 -.000.5

.lo26 -0602 - .0008.1789 -1621 a 0 0 2 5.2283 -2330 -0029

CYM-0017.0010.0012.0011.0012,0013.0009-0019.0012.0018.Ol.oe 1.9.26.1117 .3qw .1296 .0018 ,0001

03.02 20.23 1.0992 .4485 .1446 .0024 -.001022.21.1099 .SO24 .15620015.0012~ 0 4 24.21.1364 .IRIP .OOIZ -.0013

16.30.1591 .34721227 -0027 .

T EST NUMBER 19nC S F/D

- e 0 0 3 0 -5.39-.0039 3.42-.0032 9.92"0049 16.10"0069 17.82"0067 18.53"0093 16.32"0076 15.69-.0082 7.33-a0127 5.80-.0027 3.14"0066 3.66- a 0 0 4 5 2.75-a0075 8 - 4 4"0119 2-18

TEST NUMRER 198C S F/ D

-.on404.78-.0025 3.17-e0007 10.66-.0027 16.42-,OO28 18.09-.0059 17.98- e 0 0 4 8 17.00-.0081 16.05-.0100 8.25-.0097 6.76-.0072 5.27-.0081 3.26

- . O B 0 8-.OOPS-.OB39-.0056- .0060-.0105- .0054-.0099.0053- .0041-.0075-.0070

TEST NUMBER 198C S F L I D.00184.77-.0007 3.35-.00061.58 16.92

18.7918.7317.5016.118.536.114.713.312.902.502.28

T E S T NUMBER 198C S FI D

-.0119.69-.0095 9.13- . 0 1 0 0 13.40-.0124 14.13-.0178 13.11-.0179 11.59-.0145 6.94-.0153 5.69-.0153 3.34-.0131 2.80-.0146 2.45-.0154 2.21-.0170 2.08

- . o o w 14.13

H/R-427.475.522-570-618-629-630-629.630.630-629-630-630.630.630

H/R-428.475.521.569e616.630.630-630-629.630-629-629

25


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... - . - . . __

APPENDIX ARUNl l u R E R 5 L O N G I T U D I N A LS T A B I L I T Y - A X I S A ND L A T F R A L O D Y -A X I S D A TA TEST NUMRER 198M AC H O r K P A I P S f I R E T I v O E G A L P H A I D E GL CDPMRM CY H CSF L I D

RUN N l IMACH-204-204- 2 0 4- 2 0 4~ 2 0 4- 2 0 4a204- 2 0 4-204-204- 2 0 4,205.205- 2 0 5

RUN NlMACH- 2 0 4,204~ 2 0 4-204,204-204-204-204- 2 0 4- 2 0 4.PO5-205,204,205

. o n

. o o. o o. o o. o o. o o.Ol. 0 2- 0 3.04e 0 70 3- 0 4.05

-1 .99- 0 92.106 .754 .158 .2 310.3112.3314.3716 .36

20.2418.2522.2124 .29

-0611.2661.4 6 4 1,669 1

1 05681.30441 .35021.41451.15961.11791 .13871.1605

. R A S ~1.2181

L O N G I T U D I N A L T A R I L I T Y - A X ISU E T A r D E GL P H A r D E GL

- . o o.00 -1 .99 .I3630. I n- . o o 2.1 1 -2750- . o o , 4 7 6 94.14- . o o ,6776. o o 6 .1 78 .22.0575 . R T ~ B. ?. o o 10.31 1.235012.32 1 .3325. o o.04

14 .35 1 . 3 ~ 8 116 .34.4106.on 18.25.2206.03 20.25 1.1578- 0 5

22.21 1.131824 .31.1628.a4

L D N R I T U D I N A LS T A B I L I T Y - A X I SRETAsDEG ALPHAIDEGL

- .00 -2 .06 - .2725. o o. o o .0204462.03. o o 4.09. o o 6.11 -4113.Ol

.63198.21O l- . 01. 01 12 .26.1964- 0 5

14 .26.246816 .26.2876.04- 0 5 18 .24.1876- 0 4

20 .19.191922 .28.2331.064 .24.2521

. i n o r

.n5351 0 .3 7.0 6 6 4

- .I401-.1121-.0880-.OS47- .o l eo,0215

,1524.PI091440.273i?

,1766-1566.1953

. 6on

AND LATERA L nDY-cn CPM

.0250 -.I404.O269 "1153

.0409 -.0557.OS29 -a0206-0691 -0160-1460 ,1471,0886 .0599.,?OB1 12075

. o x 1 -.oaen

.26 1 1 - 2 7 9 2.3799 .1497.171n.4722 .1776,5340 -1994

. 0006,0026.OD13

.0010- n o 0 0 4.0018. o o o o-a0070- . 0060

. 0 0 1 1- 0 0 0 3- .0006-.Ob02

-.oonl

- 0 0 0 9.0010,0011- 0 0 1 3.0018. 0 0 1 1.0008- e 0 0 3 5- .0004- .0009-.0005- .0015

- a0014-.0011

.AXIS DATACUM CYM

,0024 ,0010. o m 1 .0010.0024 .0009.I3017 -0014.0015 .0016.0004 - 0 0 1 4.0002 .0012- a 0 0 2 3 -e0005.0029 .0037- .0069 .0006-a0064 -.0001.a007 - .0022,0039 - .0029- . 0 0 0 1 -.0032

- .0051 2 .46-.0086 9.R6-.0058 13.75- .0062 15.68- .0077 16 .28- .0101 14.88- .0090 13 .19-.0141 6.18-a0064 8.44- .0157 5 . 3 2-e0105 3.02- .0100 2 .61- e0104 2 .36- . o 1 1 n 2.14TEST NUMBER 198

C SF L I D- . 0 0 5 7 2.52-.0058 10 .23-.0050 14.87- .0066 16 .58-.0080 16.52-.0103 15 .29-.0090 13.95"0095 9 .1 3- .0157 6 .67"0193 5.40"0169 3.21-.0090 2 .75- .0093 2 .1 8-.0097 2.40

TEST NUMBER 198C SF L I D

-.0047 -2 .98- . 0 0 7 7 60-.00?3 2 .95-a0073 7 .9 2-.OO67 11.41-.0097 12 .96- .0111 13 .52-.0161 10.50- .0149 6.67-.0150 5 . 4 7-a0135 3 .29-.Ob82 2 .67-.0058 2 .9 3-.0107 2.38

T EST NUMRER 198C 5 F/ D

- .0047 -4 .43- .0073 1 .43-.0054 12 .98-.0076 7.87-.0067 14 .89- .0104 15 .26- .Ole4 9 .36"0114 14 .16-a0165 6.83- .0146 5 .55-.0135 3 .40- . 0 0 6 4 7.96"0109 2.65-.on91 2 .43

H / @.474,5 2 2- 5 6 8-615- 6 6 3.685-685,684.684-684.685,685,685-684

H/R,474.521,567,613.658-685.be4.685,685,685.684,684-684.684


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APPENDIX AL O N G I T U D I N A LS T A R I L I T Y - A X I S A ND L A T E RA LRODY-AXIS DATA

R E T A . n E tL P H A s D E GL CDPMRM-.00 -2.01 - .OB10 .032727790019- .00 0 4 .1*96 .n303 .21020036. o o 2.09 -3957 .0319 .16210036-6244 ,0389 .115O0018. o o.01 6.21 .n4660506079100148.30 1.0620 -0666 .0474 .0015.no 4 - 1 6

T E S T NIIMFIER 198C S F/ D

-.0070 -2.48-.00

mecanizáció aerod

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