Wind-tunnel free-flight investigation of a model of a forward-swept-wing fighter configuration

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NASA

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Technical

Paper

February 1984

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Wind-Tunnel Free-Flight

4- m e

Investigation of a ModeFE -

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of a Forward-Sw,ept-Wing-

Fighter Configuration I .

Daniel G. Murri, . . -.

Luat T. Nguyen,

and Sue B. Grafton

L O m A N COPY: RETURN TO AFWL TECHNICAL LIBRARY KIRTLAND AFB, N.M. 87117 TECH LIBRARY KAFB, N M

’- NASA

Technical

Paper

Wind-Tunnel Free-Flight

Investigation of a Model

of a Forward-Swept-Wing

-

Fighter Configuration

Daniel G. Murri, Luat T. Nguyen,

and Sue B. Grafton

Langley Research Center

Hampton, Virginia

National Aeronautics and Space Admlnlstration Scientific and Technical Information Branch SUMMARY A w i n d - t u n n e lf r e e - f l i g h ti n v e s t i g a t i o n was conducted t o s t u d y t h e s t a b i l i t y a n d c o n t r o l c h a r a c t e r i s t i c s of a model of a forward-swept-wing fighter-airplaneconfigu- r a t i o n a t highangles of attack.Otherwind-tunneltechniques employed i nt h es t u d y i n c l u d e ds t a t i c - and dynamic- ( f o r c e d - o s c i l l a t i o n )f o r c e tests, f r e e - t o - r o l l tests, and f l o w - v i s u a l i z a t i o n tests. A uniquefacet of thestudy w a s theextremelevel of s t a t i c p i t c h i n s t a b i l i t y ( i n e x c e s s of n e g a t i v e3 2 - p e r c e n ts t a t i cm a r g i n )i n h e r e n ti n theairframedesign which p r e c l u d e d f r e e - f l i g h t t e s t i n g w i t h o u t s t a b i l i t y augmenta- t i o n i n p i t c h .

The results of t h ec a p t i v e( f i x e d model)wind-tunnel tests i n d i c a t e dt h a tt h e 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 c h a r a c t e r i s t i c s of the model were determinedprimarily by the forward-swept wing and v e r t i c a l tail a t low t o moderateangles of a t t a c k and by thefuselageforehody a t thehigherangles of a t t a c k . However, component i n t e r - a c t i o n s were a l s o found t oi n f l u e n c et h e s ec h a r a c t e r i s t i c ss i g n i f i c a n t l y .

The r e s u l t s of t h ef r e e - f l i g h ts t u d y showed t h a tw i t hs t a b i l i t ya u g m e n t a t i o ni n t h ep i t c h ,r o l l , and yaw axes,the model e x h i b i t e d good dynamic s t a b i l i t yc h a r a c - t e r i s t i c s up t o 40° angle of a t t a c k , above which t h ef r e e - f l i g h tt e s t sc o u l dn o t be conducted because of lack of yaw-controleffectiveness. The p i t c hs t a b i l i t y augmen- t a t i o n system ( S A S ) w a s found t o be e f f e c t i v ei np r o v i d i n gt h er e q u i r e dl e v e l of s t a b i l i t y as long as t h e r e w a s s u f f i c i e n tp i t c hc o n t r o la v a i l a b l e . With t h er o l l SAS d e a c t i v a t e d ,t h e model exhibitedlarge-amplitude undamped r o l l o s c i l l a t i o n s (wing rock) above 25O angle of a t t a c k . The r o l l SAS actingthroughthepowerfulflaperons e f f e c t i v e l ys u p p r e s s e dt h e wing rockthroughoutthe test angle-of-attackrange.

INTRODUCTION The potentialadvantages of theforward-swept wing havebeen known f o r some time. However, a p p l i c a t i o n of theconcepthasnotbeenpracticalbecause of t h e s t r u c t u r a lw e i g h tp e n a l t yr e q u i r e dt o overcome theaeroelasticdivergenceproblem.

Recentadvances i n compositematerials and a e r o e l a s t i c t a i l o r i n g i n d i c a t e t h a t a forward-swept wing can now be b u i l t w i t h l i t t l e o r no weightpenalty.(See 1 . ) To explore this promising new technology, the Defense Advanced Research r e f .

Projects Agency (DARPA) hassponsoredthedevelopmentof a r e s e a r c ha i r p l a n e( d e s i g - nated as t h e X-29A) t od e m o n s t r a t e by f l i g h t tests t h ea p p l i c a t i o n of theforward- swept-wing conceptto a f i g h t e r - c l a s s a i r p l a n e .

A s a part of t h ee f f o r ti nd e v e l o p i n g forward-swept-wingtechnology, DARPA,, the Grumman AerospaceCorporation,andtheLangleyResearchCenterareinvolvedin a broadcooperativereseacch program toinvestigatethehigh-angle-of-attack, stall, and s p i nc h a r a c t e r i s t i c sa s s o c i a t e dw i t ht h i s advancedairplane-designconcept. A primaryobjective is to determine how t h e s e c h a r a c t e r i s t i c s are influenced by t h e unique s t a l l progression of theforward-sweptwing.Otherobjectives of t h e program are to s t u d yt h ee f f e c t s of s e v e r a l o t h e r i n t e r e s t i n g f e a t u r e s of the designinclud- ing: ( 1 1 a veryhighlevel of airframe static p i t c h i n s t a b i l i t y , (2)- a. fus.elage f o r e b o d yt a i l o r e df o rh i g h - a n g l e - o f - a t t a c kd i r e c t i o n a ls t a b i l i t y , ( 3 ) a close-coupled

p

I I I l 1 1 1 l l l l l l l

canardwithlarge-deflectioncapability, ( 4 ) full-spanflaperons, and ( 5 ) an a f t - fuselagestrakeflapforpitch-control augmentation.

The present wind-tunnel free-flight investigation was’ conducted todeterminethe s t a b i l i t y and controlcharacteristics of the X-29A configuration i n the low-speed, high-angle-of-attack ( a ) range up t o the s t a l l ( a 40° 1. The s t u d y involved wind- tunnel tests of a 0.1 6-scale model and included s t a t i c - anddynamic- (forced- oscillation)forcetests,free-to-rolltests,flow-visualizationtests, and free- flighttests.Resultsarepresented which emphasize the component effects of the variousdesignfeatures on thehigh-angle-of-attack aerodynamics and theresulting f l i g h t c h a r a c t e r i s t i c s .

SYMBOLS A l l lorLgitudina1 aerodynamic coefficientsarereferredto the stability-axis system, and alllateral-directionaldataarereferredtothe body-axissystem.

(See fig. 1.) A l l force-testdataarereferredto a moment referencecenterlocated 7 percent i n front of the wing mean aerodynamic chord.

wing span, f t Axialforce axial-forcecoefficient, q C O S Drag drag coefficient, q C O s L i f t

l i f t coefficient, -

q C O S Rolling moment rolling-moment coefficient, qmSb Pitching moment pitching-moment coefficient, X S C Normal force normal-force coefficient, q C O s Yawing moment yawing-moment coefficient, qmSb Sideforce side-forcecoefficient, % O s chord line, f t mean aerodynamic chord, f t frequency of oscillation, Hz acceleration due togravity, 32.1 52 f t / s e c moments of i n e r t i a about X , Y , and Z axes, respectively, slug-ft imaginary number, ( -1 ) ' I 2 aileron-to-rudderinterconnectgain roll-rate feedback gaintoaileron,sec roll-rate feedback gaintorudder,sec pitch-rate feedback gain to canard,sec yaw-ratefeedback gainto rudder, sec angle-of-attack feedback gainto canard reduced-frequency parameter, wb/2V or wc'/2V rollrate,rad/sec pitchrate,rad/sec f ree-stream dynamic pressure, lb/f t2 yaw rate,rad/sec wing referencearea, f t 2 time, sec time to one-halfamplitude,sec time to double amplitude,sec free-streamvelocity,ft/sec body reference axes angle of attack, deg rate of change of angle of attack,rad/sec angle of sideslip, deg rate of change of angle of sideslip,rad/sec incremental rolling-moment coefficient incremental yawing-moment coefficient incrementalside-forcecoefficient amplitude of pitchoscillation, deg amplitude of r o l l o s c i l l a t i o n , deg amplitude of yaw o s c i l l a t i o n ,d e g

-

'f, r i g h t 6 f , l e f t e f f e c t i v ea i l e r o n - d e f l e c t i o na n g l e , I deg canardincidence, positive w i t ht r a i l i n g edge down, deg c a n a r di n c i d e n c er e q u i r e df o r trim, deg f l a p e r o n - d e f l e c t i o na n g l e ,p o s i t i v ew i t ht r a i l i n ge d g e down, deg p i l o ts t i c ki n p u t ,p e r c e n t of maximum r u d d e r - d e f l e c t i o na n g l e ,p o s i t i v ew i t ht r a i l i n g edge l e f t , deg s t r a k e - f l a p - d e f l e c t i o na n g l e ,p o s i t i v ew i t ht r a i l i n g edge down, deg 2 2 damping ratio, \ I " W d / w n = sun,rad/sec angular frequency, 2rf, rad/sec damped frequency,rad/sec naturalfrequency,rad/sec S t a b i l i t y d e r i v a t i v e s : acY " ' Y - a6 B = -

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c z = - cn = - c y = -

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CI, = - cy. = -

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B ib B bb

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2v 2v 2v X, acY

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rb rb

r a -

r 3 - 2v 2v

a ‘ m acN

cm = - CN = - CA = -

qc qc 9E

g a - q a - g a -

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Abbreviations : A R I aileron-to-rudder interconnect c.g. c e n t e r of g r a v i t y S AS s t a b i l i t y a u g m e n t a t i o n s y s t e m T.E.D. t r a i l i n g edge down T.E.U. t r a i l i n g e d g e up MODELS AND TESTING TECHNIQUES Models Two models were testedto o b t a i nt h ed a t ap r e s e n t e di nt h i sr e p o r t : a f l a t - plate 16-percent-scalemodel of t h e X-29A wing planformand a 16-percent-scale model of thecomplete X-29A ( r e f e r r e d to as thebasicmodel). The geometry of the models is p r e s e n t e di nf i g u r e 2. The f l a t - p l a t e wing model w a s constructedof1/2-in-thick plywood and had a 1/4-in-radiusleadingedgeand a s h a r p ,b e v e l e dt r a i l i n g edge. The basic model w a s c o n s t r u c t e dp r i m a r i l y of molded f i b e r g l a s s and is shown i n f i g u r e 3 .

A s p i n - c h u t ec a n i s t e r that w a s t e s t e d i n three l o c a t i o n s is shown i n f i g u r e 4, and a set of 40° n o s es t r a k e st h a t w a s developedduringtesting is shown i n f i g u r e 5.

These s t r a k e s were t e s t e d i n a d d i t i o n to the standardnosestrakes which are labeled i nf i g u r e2 ( b ) .P r e s e n t e di n table I is a summary of t h ew e i g h ta n di n e r t i a s , geo- metric c h a r a c t e r i s t i c s ,a n dc o n t r o l - s u r f a c ed e f l e c t i o n s of the basic model.

The l o n g i t u d i n a lc o n t r o ls u r f a c e si n c l u d e dt h ec a n a r d , wing t r a i l i n g - e d g ef l a p - erons,and the fuselage-mountedstrakeflaps. The canardprovidedtheprimarypitch c o n t r o l of the c o n f i g u r a t i o n and had a deflectionrangefrom -6OO to 30°. The s t r a k e f l a p s are designed to augment p i t c h trim and are d e f l e c t e d to t h e f u l l trailing-edge-down position above about 18O angle of a t t a c k .I nt h ec u r r e n t i n v e s t i g a t i o n ,t h e model was t e s t e dw i t ht h es t r a k ef l a p sf i x e d a t various d e f l e c t i o n s . The f u l l - s p a nf l a p e r o n s were designedtohingesimultaneously a t t h e 75- and90-percent-chordlinesaccordingtotherelationshippresentedinfigure 6.

With t h e model inthehigh-angle-of-attackconfiguration,theflaperons and s t r a k e f l a p s were d e f l e c t e dt ot h e i rf u l l - d o w np o s i t i o n s (17.5O and 30°, r e s p e c t i v e l y ) .

However, t h ed a t ap r e s e n t e d on dynamic s t a b i l i t y d e r i v a t i v e s were measuredwiththe f l a p e r o n sd e f l e c t e dt o 2 0 ° . Unlessotherwisenoted, a l l d a t ap r e s e n t e di nt h i s r e p o r t were measuredwiththe model inthehigh-angle-of-attackconfiguration.

L a t e r a l - d i r e c t i o n a lc o n t r o l was provided by theuse of a conventionalrudder f o r yaw c o n t r o l and a s y m m e t r i c a l l yd e f l e c t e df l a p e r o n sf o rr o l lc o n t r o l .S i n c et h e model w a s testedinthehigh-angle-of-attackconfiguration,theflaperons were a t t h e i r full-downposition (17.5O) and r o l l c o n t r o l w a s provided by a s i n g l ef l a p e r o n d e f l e c t i n g up from t h a t p o s i t i o n .

Captive Wind-Tunnel Tests Extensivecaptive(fixedmodel)wind-tunneltests were conductedpriortotest- i n gt h e model i n f r e ef l i g h t . The purpose of t h e s et e s t s was t od e f i n et h e predom- i n a n t aerodynamic c h a r a c t e r i s t i c s of t h ec o n f i g u r a t i o na th i g ha n g l e s of a t t a c k and todevelop a databasetoaid i n t h ea n a l y s i s of t h ef r e e - f l i g h t test results. These tests i n c l u d e ds t a t i c - and dynamic- ( f o r c e d - o s c i l l a t i o n )f o r c e tests, f r e e - t o - r o l l t e s t s , andf l o w - v i s u a l i z a t i o n t e s t s .

The s t a t i c - and dynamic- ( f o r c e d - o s c i l l a t i o n )f o r c e tests were conducted i n t h e Langley 30- by 60-FootTunnel a t a dynamic p r e s s u r e of 10 l b / f t , correspondingto a

wing mean-chord Reynolds number of about 0.68 X 10 . I n t h e s e tests, body-axis

f o r c e s and moments were measured by using a conventionalstrain-gagebalance and were resolvedintotheappropriateaerodynamiccoefficients. The s t a t i c - f o r c e tests included component buildup tests andmeasurements of c o n t r o le f f e c t i v e n e s s and were made over an angle-of-attackrange from -90° t o 90° and an angle-of-sidesliprange from -3OO to 30°. Data a r ep r e s e n t e di nt h i sr e p o r to v e r an angle-of-attackrange from Oo t o 90° and an angle-of-sidesliprange from -5O t o 5O. To determinethe dynamic s t a b i l i t yd e r i v a t i v e s ,f o r c e d - o s c i l l a t i o nt e s t i n g was conducted i n p i t c h , r o l l , and yaw. These measurements were made a t an o s c i l l a t i o nf r e q u e n c y of 1 Hz which r e s u l t e d i n values of the reduced-frequency parameter k of 0.1 5 f o rt h er o l l - ing and yawing tests and of 0.040 f o rt h ep i t c h i n g tests.

To explorethe dynamic r o l l s t a b i l i t y of t h e model f u r t h e r , a s e r i e s of wind- t u n n e lt e s t s were performed with the model mounted on a free-to-rollapparatus.This testtechniqueallowedthe model t or o t a t ef r e e l ya b o u t its r o l l a x i s and w a s used t o r o l l o s c i l l a t i o n s a t h i g h a s s e s s t h e s u s c e p t i b i l i t y of t h ec o n f i g u r a t i o nt o undamped angles of a t t a c k (wingrock). The d a t a from these tests were p r i m a r i l yq u a l i t a t i v e ; t h a t is, t h eo b s e r v e rv i s u a l l yi d e n t i f i e dt h ep r e s e n c e and r e l a t i v e magnitude of t h e o s c i l l a t i o n s . A more d e t a i l e dd i s c u s s i o n of t h i s test technique is contained i n r e f e r e n c e 2.

L i m i t e df l o w - v i s u a l i z a t i o nt e s t s were alsoconducted by usingbothtuft- and helium-bubbletechniques. The t u f tt e c h i q u e was used t os t u d ys u r f a c ea i r f l o wo v e r t h e model,whereasthehelium-bubbletechnique was used t os t u d yp o r t i o n s of the surroundingflowfield. These tests were conducted intheLangley12-Foot Low-Speed Tunne 1.

J Free-Flight Tests The f r e e - f l i g h t test techniqueinvolvesflyingthe model u n r e s t r a i n e di nt h e open-throat test s e c t i o n of the Langley 30- by 60-Foot Tunnel. A photograph of t h e model i n f r e e f l i g h t is shown i n f i g u r e 7, and a schematicdiagram of t h e test s e t u p is shown i nf i g u r e 8. S t e a d yI gf l i g h t s were made a ta n g l e s of a t t a c k from 13O t o 40° by varyingthetunnelspeed from 96 f t / s e c (q = 11 l b / f t t o 58 f t / s e c ( g = 4 l b / f t 2 ) . The model w a s remotelycontrolled by t h r e ep i l o t s : a roll/yaw p i l o t , a p i t c hp i l o t , and a t h r u s to p e r a t o r . Pneumatic and e l e c t r i c power and con- trol s i g n a l s were supplied to t h e model through a f l e x i b l eu m b i l i c a lc o r d which was made up of wires and l i g h tp l a s t i ct u b e s . The umbilicalchordalsoincorporated a 1/8-in-steelsafetycablethatpassedthrough a p u l l e y above t h e test s e c t i o n . The s a f e t yc a b l e was used t oc a t c ht h e model when an uncontrollablemotionormechanical f a i l u r eo c c u r r e d . The e n t i r e umbilicalcord w a s k e p ts l a c kd u r i n gt h ef l i g h t s by a safety-cableoperatorusing a high-speedwinch.

The model was instrumentedwith a three-axis rate gyroscope to measure angular r a t e s and a miniaturized boom-mounted a/B vane s e n s o r( f i g .9 )t o measure angle of a t t a c k and s i d e s l i p . These d a t a ,a l o n gw i t hp i l o tc o n t r o li n p u t s and control-surface d e f l e c t i o n s , were recorded i n time-history form on s t r i p - c h a r tr e c o r d e r s . Also, qualitativeassessments of the model f l i g h t c h a r a c t e r i s t i c s w e r e taken i n t h e form of motion p i c t u r e s and p i l o t comments. The f l i g h t - c o n t r o l laws which included s t a b i l i t ya u g m e n t a t i o n i n a l lt h r e ea x e s were programmed i n t o a d i g i t a l computer which processedsensordata and p i l o tc o n t r o li n p u t st og e n e r a t e command s i g n a l st o drivethepneumaticcontrol-surfaceactuators on the model. A more d e t a i l e dd i s c u s - s i o n of t h ef r e e - fl i g h tt e s tt e c h n i q u e can be found i n r e f e r e n c e 3.

RESULTS AND DISCUSSION OF CAPTIVE TESTS S t a t i cL o n g i t u d i n a lC h a r a c t e r i s t i c s The s t a t i cl o n g i t u d i n a lc h a r a c t e r i s t i c s of theconfigurationarepresented i n f i g u r e s 10 t o 1 3 . The canard-off data of f i g u r e 10 i n d i c a t et h a tt h e wing i n i t i a l l y s t a l l e da ta b o u t 1 2 O angle of a t t a c k . The wing s t a l lp a t t e r n w a s v i s u a l i z e d by using t u f t s which showed stalledflowbeginning a t t h e wing r o o t and progressingoutward withincreasingangle of a t t a c k . The t u f t sa l s oi n d i c a t e dt h a ts m a l lr e g i o n s of attachedflow weremaintained a t the wing t i p st oa n g l e s of a t t a c kw e l l above t h a t f o r maximum l i f t ( a = 4 0 O ) . The pitching-moment data show t h a tt h ec o n f i g u r a t i o n withoutthecanardexhibits n e u t r a l t o s l i g h t l y u n s t a b l e s t a t i c s t a b i l i t y i n t h e i n t h ef r e e - f l i g h t tests ( a = 15O t o 4 0 ° ) .

angle-of-attackrangeofprimaryinterest Exceptforlargepositivecanarddeflections,addition of t h ec a n a r ds i g n i f i c a n t l y i n c r e a s e st h el e v e l of i n s t a b i l i t y€ o ra n g l e s of a t t a c k up t o 30° t o 60°,depending on canardincidence.This is followed by a stablebreak, which occurs a t an angle of a t t a c k between 3 5 O t o 60°,againdepending on canardincidence. The d a t aa l s o show thecanard t o be a veryeffectivepitch-controldeviceover a wide angle-of-attack range by v i r t u e of its largerange of t r a v e l . As shown i n f i g u r e 1 1 , the model i n thehigh-angle-of-attackconfigurationwiththec.g.at -0.07E can be trimmed through

an.angle-of-attackrange from about l o o to 7 0 ° . For theangle-of-attackrange of

p r i m a r yi n t e r e s ti nt h ef r e e - f l i g h t tests ( a = I S 0 t o 4 0 ° ) , trimmed-canard incidence varied from about 5 O t o -35O.

The d a t a of f i g u r e s 1 2 and13 show t h e e f f e c t of s t r a k e - f l a p andsymmetric f l a p - e r o nd e f l e c t i o n s . As expected, trailing-edge-down strake-flap deflections increased l i f t and,because of thelong moment arm, generatedlarge nose-down p i t c h i n g moments.

These d e v i c e sa r e shown toprovideremarkablyconstantpitch-control-momentincre- ments acrosstheentireangle-of-attackrange.Figure13 shows t h a td e f l e c t i n gt h e f l a p e r o n s to 17.5O i n c r e a s e d l i f t f o r a n g l e s of a t t a c k up to 30° andproducedsmall nose-down p i t c h i n g moments because of t h e r e l a t i v e l y s h o r t moment arm.

S t a t i c 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 C h a r a c t e r i s t i c s The static 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 c h a r a c t e r i s t i c s of t h ec o n f i g u r a t i o n are p r e s e n t e di nf i g u r e s 14 t o 22. The static s t a b i l i t yd e r i v a t i v e s C y B ~ ‘nB, and were computed using a s i d e s l i pr a n g e from -5O to 5O. The d a t a of f i g u r e 1 4

c1 B

i n d i c a t e t h a t t h e v e r t i c a l t a i l began t ol o s ee f f e c t i v e n e s si np r o v i d i n g static d i r e c t i o n a ls t a b i l i t y ( C ) a t an angle of a t t a c k between 20° t o 25O, depending on canardincidence. The loss i nv e r t i c a l - t a i le f f e c t i v e n e s s w a s due t o a blanketing e f f e c t of the t a i l i nt h e low-energy s t a l l e d wake of t h e wing.Thisresulted i n n e a r - z e r ot os l i g h t l yn e g a t i v ev a l u e s of static d i r e c t i o n a l s t a b i l i t y i n theangle- of-attackrange from 30° t o 40°. Above 40°, however, t h ed a t a show a s t r o n gr e e s t a b - lishment of s t a t i c d i r e c t i o n a l s t a b i l i t y which is obviouslynotcaused by t h e v e r t i c a lt a i l . It w a s found t h a tt h i s phenomenon was a r e s u l t of the aerodynamics associatedwiththelongslenderforebodywith a r e l a t i v e l y f l a t e l l i p t i c a l c r o s s s e c t i o n . A t highangles of a t t a c k , two s t r o n gv o r t e xs h e e t s were shed from t h e pointed nose. Under sideslip,theleewardvortex was d i s p l a c e d above the nose, whereas the windward vortex was c l o s e t o t h e s u r f a c e . ( S e e f i g . 23.) This flow patternprobablyproducedsuctionpressures on t h e windward s i d e of thenosewhich, actingthroughthelong moment arm, producedthestabilizing yawing moments. This phenomenon was observed i n past i n v e s t i g a t i o n s ( r e f s . 4 and5, f o r example ) where a similartype of forebody was tested.

Also shown i nf i g u r e 1 4 are t h e s t a t i c l a t e r a l s t a b i l i t y ( C , ) c h a r a c t e r i s t i c s B of t h ec o n f i g u r a t i o n . The data i n d i c a t e a low o ru n s t a b l ed i h e d r a le f f e c t up t o an angle of a t t a c k betweenabout So and l o o , b u tv e r ys t r o n gl a t e r a ls t a b i l i t ya tt h e higherangles of a t t a c k . A s e x p e c t e d ,t h ev e r t i c a lt a i l was s t a b i l i z i n g a t low angles of a t t a c kb e f o r el o s i n ge f f e c t i v e n e s s by a = 30°.

A comparison of t h e l a t e r a l s t a b i l i t y c h a r a c t e r i s t i c s of t h ef l a t - p l a t e wing model withthose of thecompleteconfiguration(fig.15)indicatesthattheforward- swept wing dominatedthe s t a t i c l a t e r a l s t a b i l i t y c h a r a c t e r i s t i c s of thecomplete c o n f i g u r a t i o n below about 30° angle of attack.That is, t h e wing e x h i b i t e d an unstabledihedraleffectbecause of thenegative sweep e f f e c t , when experiencing primarilyattachedflow ( a 5 loo), and a s t a b l ed i h e d r a le f f e c t , when separated vortex flow was dominant ( l o o a <, 30O). Above 30° angle of a t t a c k , component b u i l d u pt e s t si n d i c a t e dt h a t a favorableaerodynamicinteraction between theforebody and thecanard-wingcombination was r e s p o n s i b l ef o rt h eh i g hl e v e l s of s t a t i c l a t e r a l s t a b i li t y e x h i b i t e d .

The e f f e c t of c a n a r dd e f l e c t i o n on t h e s t a t i c 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 c h a r a c t e r i s t i c s of the model is p r e s e n t e di nf i g u r e 16. When deflectedtolower

s e t t i n g s ( 6 c = O o ) , thepresence of thecanardsignificantlyenhancedboth

lateral and d i r e c t i o n a ls t a b i l i t y up to about 30° angle of attack. I nt h i sa n g l e - I of-attackrange,tuft-flowphotographs(fig. 17) i n d i c a t e dt h a tf o r l o w d e f l e c t i o n s , thecanardimprovedtheflowfield a t t h ev e r t i c a l t a i l . A t t h eh i g h e rd e f l e c t i o n s (6c = -60°), however, t h ec a n a r dc o n t r i b u t e dt ot h eb l a n k e t i n ge f f e c t of t h e wing, thuscausing a d e g r a d a t i o ni nb o t h lateral a n dd i r e c t i o n a ls t a b i l i t y . Between 30° and 45O angle of a t t a c k , t h e d a t a of f i g u r e 16 show t h a tc a n a r dd e f l e c t i o np r i m a r i l y a f f e c t e dd i r e c t i o n a ls t a b i l i t y .I nt h ea n g l e - o f - a t t a c kr a n g e from 30° t o 45O, t h e lowercanarddeflections(6c = Oo and -25O) r e s u l t e d i n a d e g r a d a t i o ni nd i r e c t i o n a l s t a b i l i t y , as compared withthecanard-offconfiguration.

P r e s e n t e d i n f i g u r e s 18 and 19 are t h e e f f e c t s of f l a p e r o na n ds t r a k e - f l a p d e f l e c t i o n s ,r e s p e c t i v e l y , on t h e s t a t i c 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 yc h a r a c t e r - istics. Symmetricflaperondeflection(fig. 18) is shown t o have a minimal e f f e c t on both lateral and d i r e c t i o n a ls t a b i l i t y . On t h eo t h e r hand, t h ed a t a of f i g u r e 19 i n d i c a t e t h a t s t r a k e - f l a p d e f l e c t i o n s h a v e a s i g n i f i c a n t e f f e c t on d i r e c t i o n a l sta- b i l i t y up t oa b o u t 55O angleofattack. B e l o w 25O angleofattack,trailing-edge-up s t r a k e - f l a p d e f l e c t i o n s d e g r a d e d d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s , p r o b a b l y as a r e s u l t of a d v e r s ei n t e r f e r e n c ew i t ht h ev e r t i c a l - t a i lf l o wf i e l d . Between 25O and 55O angle of a t t a c k , however,trailing-edge-updeflectionsimproveddirectional s t a b i l i t y .

The s t a t i c s t a b i l i t y c h a r a c t e r i s t i c s of theconfigurationwiththespin-chute c a n i s t e r are p r e s e n t e di nf i g u r e 20. The d a t ai n d i c a t et h a tt h es p i n - c h u t ec a n i s t e r has a minor e f f e c t on 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 below 45O angle of a t t a c k f o r a l l t h r e el o c a t i o n st e s t e d .

Presented i n f i g u r e 21 a r et h ee f f e c t s of addingthe 40° nose s t r a k e s on the s t a t i c 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 . As w i l l be discussed later, t h e s es t r a k e s were developed t o i n c r e a s e r o l l damping i nt h er e g i o n of maximum l i f t .F i g u r e 21 i n d i c a t e s , however, t h a tt h e yh a v e a v e r yd e t r i m e n t a le f f e c t on static d i r e c t i o n a l s t a b i l i t y . With t h e 40° nosestrakesadded, a r e d u c t i o ni nd i r e c t i o n a ls t a b i l i t y w a s measured above 5 O angle of a t t a c k and negative values of C were measured a t “ B anglesofattackabove 2 0 ° . This would i n d i c a t et h a tt h e 40° n o s es t r a k e si n h i b i t e d theformation of t h ed i r e c t i o n a l l ys t a b i l i z i n gv o r t e xf l o wf i e l dn o r m a l l yp r o d u c e d by theforebody. A l s o , a reduction i n l a t e r a l s t a b i l i t y above S O 0 angle of attacksug- g e s t st h a tt h e 40° nosestrakesdegradethefavorableflow-fieldinteractionbetween theforebodyandthecanard-wingcombination.

The e f f e c t of t h e boom-mounted a/$ vane s e n s o r( f i g . 9 ) used i nt h ef r e e - f l i g h t tests on t h e s t a t i c 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 is p r e s e n t e di nf i g u r e 22.

The d a t a i n d i c a t e t h a t t h e p r i m a r y e f f e c t of thesensorandattachment boom w a s t o d e g r a d e s l i g h t l y t h e d i r e c t i o n a l s t a b i l i t y above 25O angle of a t t a c k .

G e n e r a l l y ,t h er e s u l t s of t h e s t a t i c wind-tunnel tests i n d i c a t et h a tt h e 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 c h a r a c t e r i s t i c s of theconfiguration were not domi- nated by any one configurationfeaturethroughouttheangle-of-attackrange. The d i r e c t i o n a l s t a b i l i t y w a s found t o be s t r o n g l yi n f l u e n c e d a t t h e loweranglesof a t t a c k by t h e v e r t i c a l t a i l andcanards,whereastheforebodydesign w a s found t o dominate a t highanglesofattack. The forward-swept wing w a s p r i m a r i l yr e s p o n s i b l e f o r t h e lateral s t a b i l i t y c h a r a c t e r i s t i c s up t oa b o u t 30° angleofattack;however, t h ec h a r a c t e r i s t i c sa b o v e 30° angleofattackappearedto be s t r o n g l yi n f l u e n c e d by i n t e r a c t i o n between theforebodyandthecanard-wingflowfields.

L a t e r a l - D i r e c t i o n a l C o n t r o l C h a r a c t e r i s t i c s The r e s u l t s of t e s t sc o n d u c t e dt od e t e r m i n et h ee f f e c t i v e n e s s of t h e lateral- d i r e c t i o n a lc o n t r o ls u r f a c e sa r ep r e s e n t e di nf i g u r e s 24 t o 27. The d a t aa r ei nt h e formof incrementalforces and moments producedby a i l e r o n or rudderdeflections.

Because t h e l a t e r a l - d i r e c t i o n a l c o n t r o l e f f e c t i v e n e s s was found t o be somewhat depen- d e n t on canardincidence,data are p r e s e n t e df o rs e v e r a lc a n a r dd e f l e c t i o n s . Shown i n f i g u r e 24 a r e t h e e f f e c t s of f u l l and i n t e r m e d i a t e a i l e r o n ( d i f f e r e n t i a l f l a p e r o n ) d e f l e c t i o n s . As expected,thedata show t h a t a t low angles of a t t a c kt h ef u l l - s p a n flaperonsprovidedlarge amounts of r o l lc o n t r o l ; however, as wing s t a l l developed a t about a = loo, a i l e r o ne f f e c t i v e n e s s w a s sharply reduced. Nonetheless, for the full d e f l e c t i o n a s i g n i f i c a n t and n e a r l yc o n s t a n tl e v e l of r o l l - c o n t r o le f f e c t i v e n e s s w a s maintained a t an angle of a t t a c k from 25O up t o 60°. Duringtuft-flowstudies, regions of attachedflow were observed a t t h e wing t i p s a t angles of a t t a c k as high as 60°. Thus, it is apparentlythenature of the forward-swept-wing s t a l l progres- s i o nt oa l l o wt h ea i l e r o n st om a i n t a i ne f f e c t i v e n e s s up t o veryhighangles of a t t a c k . Comparison of the two sets of d e f l e c t i o nd a t ai n d i c a t et h a tt h ea i l e r o n a = 30°. Above 30°, e f f e c t i v e n e s s was f a i r l yl i n e a rw i t hd e f l e c t i o n up t oa b o u t however, t h ed a t a show t h a t ,p r o p o r t i o n a t e l y ,t h es m a l l e rd e f l e c t i o n w a s much l e s s e f f e c t i v et h a nt h el a r g e rd e f l e c t i o n . The yawing-moment r e s u l t s show t h a tt h ea i l e - ronsproducedadverse yaw above 15O angle of a t t a c k ; however, t h el e v e l of t h e adverse yawing moments remained relativelysmallthroughoutthehigh-angle-of-attack range.

A s would be e x p e c t e d ,t h ee f f e c t of c a n a r dd e f l e c t i o n on a i l e r o ne f f e c t i v e n e s s w a s found t o result from changes i nt h e wing f l o wf i e l d . Comparison of d a t af o r 6 = O o t o canard-offdatainfigure 25 i n d i c a t e st h a tt h ec a n a r ds i g n i f i c a n t l y C enhancestheaileroneffectivenessintheangle-of-attackrange from loo to 60°.

T h i se f f e c t w a s probablyduetothecanardimprovingtheflowovertheinboardpor- t i o n of the wing. The d a t af o r g C = -25O and -6OO i n d i c a t et h a tt h eh i g h e rc a n a r d d e f l e c t i o n s r e d u c e d o r e l i m i n a t e d t h i s b e n e f i c i a l e f f e c t e i t h e r by d i r e c t l y b l a n k e t - ingtheinboardsection of t h e wing w i t hs t a l l e df l o w from thecanard or by simply promoting wing s t a l l by i n c r e a s i n g upwash.

Presented i n f i g u r e 26 aretheincrementalforces and moments due t o one-half and f u l l rudderdeflections. The d a t a show t h a tt h er u d d e re f f e c t i v e n e s sd e t e r i o - r a t e d as wing stall developed, and it reached'very low values by approximately 40° angle of a t t a c k . A s d i s c u s s e dp r e v i o u s l y ,t h i se f f e c t w a s p r i m a r i l y due t ot h e b l a n k e t i n g of t h ev e r t i c a lt a i li nt h e low-energy s t a l l e d wake of the wing. The e f f e c t of thecanard on rudder power ( f i g . 27) a p p e a r st o be v e r yc o n s i s t e n tw i t h thecanardeffects on d i r e c t i o n a ls t a b i l i t yd i s c u s s e de a r l i e r . The canard a tz e r o i n c i d e n c e s i g n i f i c a n t l y enhancedruddereffectiveness by suppressing wing stall and therefore improving the flow a t t h ev e r t i c a l tail. Unloading (downloading) the c a n a r dr e d u c e dt h i sb e n e f i c i a le f f e c t .

Dynamic S t a b i l i t y D e r i v a t i v e s The e f f e c t s of thecanard on the dynamic l o n g i t u d i n a l s t a b i l i t y d e r i v a t i v e s o b t a i n e dd u r i n gp i t c h i n g - o s c i l l a t i o n tests a r e shown i nf i g u r e 28. The canard-off c o n f i g u r a t i o ne x h i b i t e ds t a b l ep i t c h - d a m p i n gc h a r a c t e r i s t i c s up t o a = 70°, which suggeststhatthefuselagestrakesprovided much of t h e damping e x h i b i t e d by t h e completeconfiguration.Addition of thecanard had t h es t r o n g e s te f f e c t a t low a n g l e s of a t t a c k ( a < 15O) and less s i g n i f i c a n te f f e c t s a t higherangles of a t t a c k .

The dynamic 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 yd e r i v a t i v e so b t a i n e dd u r i n gf o r c e d - o s c i l l a t i o n tests i n r o l l are p r e s e n t e di nf i g u r e s 29 t o 3 1 . The d a t ao ff i g u r e 29 show that the r o l l damping decreased above a = l o o withtheonset of wing stall suchthatunstable damping w a s encounteredabove 20° angle of a t t a c k . With the

canard removed, t h ed a t a show highlyunstablevalues of C, + C,. s i n a t h a t . were

P B measured intheangle-of-attackrange from 30° t o 45O, where theslenderforebody producedthestrongvortexflowfield that dominated static s t a b i l i t y c h a r a c t e r i s t i c s as discussed earlier. A s expectedbecause of delayedwing-tip s t a l l , t h ed a t a of f i g u r e 30 show stable roll damping fortheflat-plateforward-swept wing i n the angle-of-attackrangefrom 30° to 40°. The severe loss of roll damping- e x h i b i t e d by the complete configuration above a = 20° i s , t h e r e f o r e ,n o ta t t r i b u t a b l et o the wing planform. Comparison of these r e s u l t sw i t ht h o s e of p a s ti n v e s t i g a t i o n s (see, f o r example, r e f . 4 ) of otherconfigurationshaving a similar forebodystrongly suggeststhattheforebodyaerodynamicsdominatedtheroll damping at highangles of a t t a c k andcausedtheunstablecharacteristics. The d a t a of f i g u r e 29 i n d i c a t et h a t a d d i t i o n of thecanard a t appropriatedeflectionscanreducetheseverityofthe i n s t a b i l i t y b u t n o t e l i m i n a t e it.

I n anattempttofindsolutionstothisroll-dampingproblem, a number of fore- body s t r a k e s w e r e i n v e s t i g a t e dw i t ht h ei d e at h a tt h e yc o u l ds i g n i f i c a n t l y a l t e r the v o r t e xf l o wf i e l ds h e d from theforebody and t h e r e f o r e improve r o l l damping. The designthatprovidedthe most improvement was t h e 40° nose s t r a k ed i s c u s s e d earlier.

(Seefig. 5 . ) As shown i nf i g u r e 31 , this s t r a k e enhanced r o l l damping i nt h ea n g l e - of-attackrange from 20° t o 40°; n o n e t h e l e s s ,o v e r a l l damping remained very low i n t h i s angle-of -attack range.

The poorroll-dampingcharacteristicscoupledwithstrong s t a t i c l a t e r a l sta- b i l i t y and a h i g hi n e r t i ar a t i o of yaw t or o l l ( I /I I O ) would be e x p e c t e dt o

z x

make t h es t u d yc o n f i g u r a t i o ns u s c e p t i b l et o undamped r o l l o s c i l l a t i o n s , r e f e r r e d t o as wing rock, above 20° angle of a t t a c k . To e x p l o r et h i s phenomenon f u r t h e r ,f r e e - t o - r o l l tests were conductedand a summary oftheresultsobtained is p r e s e n t e di n t a b l e 11. The r o l lm o t i o n s of thecompleteconfiguration were found t o be l i g h t l y damped by a = 20°. A s angle of a t t a c k was f u r t h e ri n c r e a s e d , limit cycle (undamped) wing rock was observed which began a t a = 25O and reached large amplitudes ( A + > k25O) above a = 35O. A l s o , removal of t h ec a n a r d ,v e r t i c a l tail, or wing i n d i v i d u a l l yd i dn o tf u n d a m e n t a l l ya l t e rt h e wing-rock c h a r a c t e r i s t i c s . These r e s u l t s providefurtherevidencethattheaerodynamics of thefuselage were t h e source of t h eu n s t a b l er o l l damping e x h i b i t e d by the test configuration. Tests with t h e 40° nose s t r a k e s showed thatalthoughthestrakereducedtheamplitude of t h e wing rock, it d i dn o te l i m i n a t e it.

The r e s u l t s of t h ef o r c e d - o s c i l l a t i o n tests i n yaw are p r e s e n t e di nf i g u r e s 32 and 33. Figure 32 shows t h a tw i t ht h ec a n a r dd e f l e c t e dt o -25O, t h ev e r t i c a l t a i l provided a s t a b i l i z i n gi n c r e m e n tt ot h e yaw-damping d e r i v a t i v e C - C . cos a "r "B throughouttheangle-of-attackrangetested. The d a t a of f i g u r e 33 show t h a tt h e canard had t h e g r e a t e s t e f f e c t on yaw dampingbetween angles of a t t a c k of about 25O and 50°. In thisangle-of-attzckrange,thebest yaw-damping c h a r a c t e r i s t i c s were obtained a t themoderatecanarddeflections ( 6 = -25O), probably as a r e s u l t of C i n c r e a s e d v e r t i c a l - t a i l e f f e c t i v e n e s s .

1 1 FREE-FLIGHT-MODEL CONTROL LAWS Based on the static anddynamicwind-tunnel d a t ad i s c u s s e d earlier, it was obvi- ous that the model would be u n f l y a b l e w i t h o u t s t a b i l i t y a u g m e n t a t i o n i n t h e p i t c h a x i sa n at h a t it c o u l db e n e f i tg r e a t l y from augmentationin r o l l and yaw. A s a r e s u l t , a three-axisstabilityaugmentationsystem ( S A S ) w a s developed. The SAS gains were d e f i n e df o rt h ef u l l - s c a l ea i r p l a n e by u s i n gs t a n d a r dl i n e a r - a n a l y s i s techniques, and thenthey were dynamicallyscaledforimplementationinthe model controlsystem.Althoughthehandling-qualityspecifications of MIL-F-8785B(ASG) ( r e f . 6 ) do n o ta p p l yd i r e c t l yt o 'h i g h - a n g l e - o f - a t t a c kf l i g h tc o n d i t i o n s , these s p e c i f i c a t i o n s were used i n t h e SAS designprocess as rough g u i d e l i n e sf o rd e f i n i n g desiredfrequency and damping c h a r a c t e r i s t i c s . For example, f i g u r e 34 i l l u s t r a t e s theprocedureusedtodefinethepitch-axisaugmentationforIgflight a t 30° angle ofattack. Note t h a tt h er e s u l t sp r e s e n t e dc o r r e s p o n d to t h ef u l l - s c a l e airplane c h a r a c t e r i s t i c s . Because of t h eh i g h l yu n s t a b l e s t a t i c characteristics, t h e basic airframeshort-period mode was aperiodicwith a time to doubleamplitude of 0.85 sec.

A s shown i nt h er o o t - l o c u sp l o t , a SAS usingangle of a t t a c k and pitch-ratefeedbacks canprovide the desiredshort-periodfrequency and damping c h a r a c t e r i s t i c s . The feedback gains shown ( K a = -2.0 and = -1 .O sec) were s c a l e d andused i n the b a s e l i n ep i t c h SAS t h r o u g h o u tt h ef r e e -l i g h t tests. A d i s c u s s i o n on t h el i n e a r - 2 analysistechnique and t h e u s e of theroot-locusplotcan be found i nr e f e r e n c e 7.

A correspondingexampleforaugmentation of the Dutch r o l l mode is shown i n f i g u r e 35. A t a = 30°, the basic airframe Dutch r o l l mode was undamped, p r i m a r i l y because of t h eu n s t a b l er o l l - d a m p i n gc h a r a c t e r i s t i c sd i s c u s s e d earlier. The root- l o c u sp l o ti n d i c a t e st h a tt h ed e s i r e d Dutch r o l lc h a r a c t e r i s t i c sc o u l d be obtained throughroll-dampingaugmentation by d e f l e c t i n gt h ea i l e r o n sw i t hr o l l - r a t ef e e d b a c k a t a f a i r l y h i g h g a i n .

The three-axisstabilityaugmentationsystemdevelopedforthefree-flight tests is summarized i nf i g u r e 36 and is giveninthefollowing table: I na d d i t i o nt ot h ep i t c h and r o l lc h a n n e l sd i s c u s s e dp r e v i o u s l y , a yaw SAS com- p r i s i n g yaw and r o l l - r a t ef e e d b a c k s w a s a l s o implemented. The yaw-rate feed- back ( K r = 2.0 sec) augmented yaw damping, and t h er o l l - r a t ef e e d b a c k (Kpr = -0.4 sec) enhanced roll coordination. A n aileron-to-rudderinterconnect ( A R I ) w a s a l s o implemented t o improve r o l l c o o r d i n a t i o n a t highanglesofattack by causing the model t o r o l l more nearlyabout its v e l o c i t yv e c t o r . The e f f e c t of the A R I on 1 2 Positivevalues of t h i sp a r a m e t e ri n d i c a t e a normal r o l lr e s p o n s e , and negative v a l u e si n d i c a t e a reversedresponse. The d a t a of f i g u r e 37 i n d i c a t et h a tw i t h o u tt h e ARI, t h e model would e x h i b i t a reversedresponseto a l a t e r a l - c o n t r o li n p u t above about 30° angle of attackbecause of a i l e r o na d v e r s e yaw. With the a d d i t i o n of the A R I (KARI = 2.31, h o w e v e r ,t h ed a t ai n d i c a t et h a t normalresponsecould be achieved up t o 40° angle of a t t a c k . The r e d u c t i o ni nt h e ARI-on LCDP values above a = 30° was due to thecombination of a i l e r o na d v e r s e yaw and loss i nr u d d e re f f e c t i v e n e s s .

Based on t h e s e results, it would be e x p e c t e dt h a tt h el a t e r a l - d i r e c t i o n a lc o n t r o l of the model i n yaw and r o l l i n f r e e f l i g h t would degraderapidly above 30° angle of a t t a c k and be marginal by a = 40°.

It should be notedthattheaforementionedcontrol laws were developedto meet onlytherequirementsdictated by t h er e s t r i c t e dt e s tc o n d i t i o n s of t h ef r e e - f l i g h t t e s tt e c h n i q u e - t h a t is, r e l a t i v e l ys m a l l motion perturbationsabout a l g wings- l e v e lf l i g h tc o n d i t i o n . As a r e s u l t ,t h ec o n t r o l laws d e r i v e da r ef a i r l ys i m p l e .

The design of c o n t r o l laws forthefullhigh-angle-of-attackenvelope of a highly maneuverable airplanesuchasthe X-29A r e q u i r e st h a t many o t h e rf a c t o r s be addressed which w i l l inherentlycomplicatethefinaldesign. A d e t a i l e dd i s c u s s i o n of some of i n thedesign of high-angle-of-attackcontrol laws is presented thefactorsinvolved i n reference 8.

RESULTS AND DISCUSSION OF FREE-FLIGHT TESTS BaselineConfiguration The primaryconfigurationinvestigated i n t h ef r e e - f l i g h tt e s t s w i l l be r e f e r r e dt oa st h eb a s e l i n ec o n f i g u r a t i o n and is definedasfollows:

-

c.g. = -0.12c 6 f = 17.5O 6 s = 30° SAS on; A R I on L o n g i t u d i n a lc h a r a c t e r i s t i c s . - The l o n g i t u d i n a l s t a b i l i t y and controlcharac- t e r i s t i c s of t h eb a s e l i n ec o n f i g u r a t i o n were found t o be good a t a n g l e s of a t t a c k from 20° up t ot h e maximum a t t a i n e di n the. t e s t s ( 4 0 O ) . The model was easy t o f l y and a p p e a r e dt ot h ep i l o tt o have good s t a t i c s t a b i l i t y and damping c h a r a c t e r i s t i c s .

These r e s u l t s v e r i f i e d t h a t t h e p i t c h SAS w a s e f f e c t i v e i n masking thehighlevel of i n h e r e n ti n s t a b i l i t y of t h eb a s i ca i r f r a m e . The o n l yi n d i c a t i o n st ot h ep i l o tt h a t he w a s f l y i n g a h i g h l yu n s t a b l ec o n f i g u r a t i o n were the v a r i a t i o n of trimmed-canard d e f l e c t i o n w i t h angle of a t t a c k and theunusuallevel of c a n a r da c t i v i t ya p p a r e n t a t a l lf l i g h tc o n d i t i o n s . It w a s o b s e r v e dt h a td u r i n gf l i g h t ,t h ec a n a r d s w e r e 1 3 c o n s t a n t l yu n d e r g o i n gr e l a t i v e l y small d e f l e c t i o n sa b o u t the trim s e t t i n g ,e v e n duringperiodswithout p i l o t p i t c hi n p u t s .T h i sc h a r a c t e r i s t i cr e s u l t e d from t h e f a c t t h a tt h eh i g h - g a i np i t c h SAS was t o t a l l yr e s p o n s i b l e for t h ea i r p l a n ep i t c h s t a b i l i t y and t h e r e f o r e had to respond t o a l l disturbances,such as thosecausedby turbulence or coupling from theroll/yawaxes.

Despite theconstantmotion, however, thecanard-deflection ( s l e w ) rates o b s e r v e di nt h ef r e e - f l i g h t tests were nothigh. Maximum slew rates didnotexceed 100°/sec (40°/sec f u l l scale). It should be n o t e d ,h o w e v e r ,t h a tt h el gf l i g h t conditionofthe tests didnotprovidethe most severe t e s t ofcanard slew-rate requirements. It would be e x p e c t e dt h a t much higher rates would be neededduring rapid,large-amplitudemaneuverssuchasthoseencounteredduringclose-in air-to- a i r combat.

An u n d e s i r a b l ec h a r a c t e r i s t i cn o t e di nt h ef l i g h t tests w a s a lackofadequate nose-up p i t c hc o n t r o l a t a n g l e s of a t t a c k below 20°. The trimmed-canard s e t t i n gf o r t h e s ec o n d i t i o n s w a s approximatelyzeroand, as shown i nf i g u r e1 0 ,v e r y l i t t l e nose- be generated by t r y i n g to i n c r e a s e l i f t on thecanard(load up p i t c h i n g moment could t h ec a n a r df u r t h e r )s i n c et h a t would only promote canard s t a l l . A possible s o l u t i o n t o t h i s problem would be to o b t a i na d d i t i o n a l nose-up c o n t r o l moment by making t h e s t r a k ef l a pa na c t i v ep i t c h - c o n t r o ld e v i c e . As discussed earlier, t h es t r a k ef l a p s provide a n e a r l yc o n s t a n tl e v e l of controleffectivenessthroughouttheangle-of- attackrange.Furtherimprovementscouldalso be obtained by modifyingthestrake- f l a pd e f l e c t i o ns c h e d u l e so t h a t f u l l nose-down d e f l e c t i o n is n o tr e a c h e du n t i l a is approximately30°.

I nc l o s i n gt h ed i s c u s s i o n of l o n g i t u d i n a lc h a r a c t e r i s t i c s , it i s a p p r o p r i a t et o notethattheaccomplishment of t h ef r e e - f l i g h t tests of t h e X-29A model r e p r e s e n t s s e v e r a l firsts for t h es t a l l / s p i nr e s e a r c h program a t Langley. This work i s the f i r s t time t h a t a dynamicallyscaled,remotelypilotedfree-flight model hasbeen successfullyflown up throughthe s t a l l withsuch a high level ofairframepitch i n s t a b i l i t y ( i n excess of -32 p e r c e n t ) . It is also t h e f i r s t time t h a ts u c h a high l e v e l ofdependencehasbeenplaced on theautomaticstabilityaugmentationsystem f o rs a t i s f a c t o r yf l i g h tc h a r a c t e r i s t i c s . The r e s u l t s of the tests i n d i c a t et h a t a i r p l a n ed e s i g n is f e a s i b l ef o rv e r yh i g hl e v e l s of a i r f r a m ep i t c hi n s t a b i l i t y as long as t h e r e is adequatepitch-control power and sensorinformationtoprovidethe necessarylevelsofstabilityaugmentation.

L a t e r a l - d i r e c t i o n a lc h a r a c t e r i s t i c s . - The p o s s i b i l i t y of d i r e c t i o n a ld i v e r g e n c e is normallyexaminedby means ofthedynamic d u r i n g f l i g h t a t highanglesofattack

. (See ref. 9.) Negative values of this

d i r e c t i o n a ls t a b i l i t y parameter Cn B , d P parameter indicate the existence of a d i r e c t i o n a ld i v e r g e n c e .V a l u e s of Cn

B , dyn

c a l c u l a t e d from t h es t a t i cw i n d - t u n n e ld a t ad i s c u s s e d e a r l i e r are p r e s e n t e di nf i g - u r e 38. Because canard s e t t i n g had a s i g n i f i c a n te f f e c t on C , t h e Cn "B B, dYn values were computed based on t h ec a n a r dd e f l e c t i o nr e q u i r e df o r trim a t eachangle of attack. It is s e e n t h a t Cn remained p o s i t i v e t h r o u g h o u t t h e test angle-of-

B , dYn

a t t a c kr a n g e ,t h u si n d i c a t i n gt h a t a d i r e c t i o n a ld i v e r g e n c e would n o t be encountered during the f r e e - f l i g h t tests.

The 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 and c o n t r o lc h a r a c t e r i s t i c so b s e r v e df o rt h e b a s e l i n ec o n f i g u r a t i o n are summarized i nf i g u r e 39. B e l o w about 25O angleofattack, t h e model w a s e a s yt of l y ,t h u se x h i b i t i n g good s t a b i l i t y c h a r a c t e r i s t i c s and high c o n t r o le f f e c t i v e n e s s . It w a s noted, however, t h a tt h e yaw motions were somewhat unsteady, as i f the model w a s alwaysbeingexcited by small d i s t u r b a n c e si n yaw.

Thesedisturbancescouldhaveresulted from s e v e r a lf a c t o r ss u c h as tunnelturbu- lence; however, it is f e l t t h a t t h e c o n s t a n t c a n a r d motiondiscussed earlier may a l s o have contributedbecause of t h es t r o n gi n f l u e n c e of thecanard on t h e v e r t i c a l - t a i l f l o wf i e l d .I n the angle-of-attackrange from 25O t o 35O, mild nose wandering was encountered, probably as a r e s u l t of loss of s t a t i c d i r e c t i o n a ls t a b i l i t y .( S e e f i g . 16.) The nose wandering, coupled with degrading rudder effectiveness (see f i g . 2 7 ) , i n c r e a s e dt h ep i l o t work l o a dr e q u i r e dt of l yt h e model. Nonetheless, theflightmotionsremained well-damped and t h e r e w a s still s u f f i c i e n t lateral- d i r e c t i o n a lc o n t r o le f f e c t i v e n e s s . Above 35O angle of a t t a c k , however, therudder e f f e c t i v e n e s s had degradedto a p o i n t where c o n t r o l of t h e model became marginaland p i l o t work load w a s veryhigh.Generally,near 40° angle of a t t a c k ,c o n t r o l of t h e model w a s l o s t i n a slow l a t e r a l t r a n s l a t i o no u t of the j e t a g a i n s t f u l l c o r r e c t i v e c o n t r o l . To e l i m i n a t et h ep o s s i b i l i t yt h a tt h ea i l e r o na d v e r s e yaw w a s causingthe d e p a r t u r e ,f l i g h t s were made w i t ht h ea i l e r o nd e a c t i v a t e d so t h a tt h er u d d e r was the o n l yl a t e r a l - d i r e c t i o n a lc o n t r o l . Loss of c o n t r o l w a s alsoexperiencednear 40° angle of a t t a c k ,t h u sf u r t h e rc o n f i r m i n gt h a t loss of r u d d e re f f e c t i v e n e s s w a s the primarycause of thedivergence.

E f f e c t of S t a b i l i t y Augmentation Systems P i t c ha x i s . - The e f f e c t of thepitchstabilityaugmentationsystem (SAS) on t h el o n g i t u d i n a lf l y i n gc h a r a c t e r i s t i c s of the model w a s evaluated by successively Presented i nf i g u r e 40 reducingtheangle-of-attack and pitch-ratefeedbackgains.

a r e time h i s t o r i e s of f l i g h t s made withthebaseline SAS and withthe SAS gains reduced by 50 percent. W i t h t h eb a s e l i n e SAS, it is s e e nt h a tt h el o n g i t u d i n a l motions of the model were well-behaved. Small p e r t u r b a t i o n si na n g l e of a t t a c k and p i t c h rate were experienced, and t h e p i t c h p i l o t was r e q u i r e dt o make onlyvery small inputstomaintainthe model i nt h ed e s i r e da r e a of t h et e s ts e c t i o n . With the SAS gainsreduced by 50 percent, however, the time h i s t o r i e s show significantlydegraded c h a r a c t e r i s t i c s .D e s p i t e numerous r a p i dp i l o ti n p u t s ,p i t c hc o n t r o l of the model was very poor as i n d i c a t e d by t h el a r g ee x c u r s i o n si na n g l e of a t t a c k . It w a s found t h a t the model could be flownwithpitch SAS gains less than 50 percent of t h eb a s e l i n e values;however,the p i l o t work load w a s extremelyhigh and the model pitchexcur- s i o n s were t o ol a r g et oa l l o we f f e c t i v ee v a l u a t i o n of its 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 s .

R o l l axis.- The e f f e c t of t h e r o l l SAS w a s s t u d i e d by making f l i g h t s w i t h reduced r o l l SAS gains andwiththeroll SAS deactivated.Figure 41 p r e s e n t s a sum- mary of the model l a t e r a l - d i r e c t i o n a lc h a r a c t e r i s t i c s when flownwithouttheroll SAS. Below 20° angle of a t t a c k ,t h e model e x h i b i t e d good f l y i n gc h a r a c t e r i s t i c s , although damping i n r o l l w a s n o t i c e a b l y lowerthan t h a t of thebaselineconfigura- t i o n . Above 20° angle of attack,small-amplitude, undamped r o l lo s c i l l a t i o n s were encountered which increased p i l o t work load.Control of the model became marginal above about 25O angle of a t t a c k , and loss of c o n t r o lu s u a l l yo c c u r r e d by 30° angle of attackbecause of large-amplitude wing rock.

The rapidbuildup of t h e wing rock w a s s t u d i e d by making f l i g h t s i n which t h e r o l l SAS w a s d e a c t i v a t e df o rs h o r tp e r i o d s of time. D u r i n gt h e s ep e r i o d s ,r o l lp i l o t i n p u t s were a l s o k e p t t o a minimum t oa v o i da f f e c t i n gt h e damping c h a r a c t e r i s t i c s o f t h e model. P r e s e n t e di nf i g u r e 42 are time h i s t o r i e s measured during one of these f l i g h t s a t a trim angle of a t t a c k of about 25O. S h o r t l ya f t e rd e a c t i v a t i o n of t h e r o l l SAS (t = 7 s e c ) , wing rock was encounteredwiththemotionsrapidlyincreasing in amplitude. Comparison of t h e p and r traces show t h a t t h e o s c i l l a t i o n s were p r i m a r i l ya b o u tt h er o l la x i s and t h e r e f o r e were very similar tothoseobservedin t h e f r e e - t o - r o l l tests. At t 18 s e c , r e a c t i v a t i o n of t h er o l l SAS was necessary to prevent loss of c o n t r o l .of t h e model. It is seen that t h e r o l l SAS veryquickly suppressedthe wing-rock motion. Flightsconductedwithreducedroll SAS g a i n s showed t h a t s a t i s f a c t o r y f l y i n g c h a r a c t e r i s t i c s c o u l d n o t be achievedwith more than a 50-percentreduction of t h eb a s e l i n er o l l SAS gain.

The aforementioned results confirmedthepredictions of t h ef o r c e d - o s c i l l a t i o n and f r e e - t o - r o l l t e s t s and v e r i f i e d t h a t a h i g h - g a i nr o l l SAS couldsuppressthe wing rockinherentintheairframe.

Yaw axis.- As e x p e c t e d ,d e a c t i v a t i n gt h e yaw SAS s i g n i f i c a n t l yr e d u c e dt h e yaw it more d i f f i c u l tt of l y . The unsteadiness of t h e damping of the model which made yaw motions were magnified and p i l o t work loadincreasedaccordingly. However , eliminatingthe yaw SAS d i dn o ts i g n i f i c a n t l y alter t h e wing-rock c h a r a c t e r i s t i c s o r t h e a b i l i t y t o f l y t h e model up t o t h e maximum l i f t r e g i o nb e f o r ec o n t r o l was l o s t because of lack of ruddereffectiveness.

Conf i g u r a t i o n E f f e c t s I na d d i t i o nt ot h eb a s e l i n ec o n f i g u r a t i o n ,f l i g h t s were a l s o made t o determine t h ee f f e c t s of t h es p i n - c h u t ec a n i s t e r( a tt h r e el o c a t i o n s ) , the 40° nose s t r a k e s , t h e c e n t e r - o f - g r a v i t y l o c a t i o n , and t h e s t r a k e - f l a p d e f l e c t i o n .

The f l i g h t c h a r a c t e r i s t i c s of t h e model withtheaddition of thespin-chute c a n i s t e r i n any of t h et h r e el o c a t i o n s were found t o be very similar tothecharac- t e r i s t i c s of t h eb a s e l i n ec o n f i g u r a t i o n .T h i sr e s u l ta g r e e sv e r yw e l lw i t ht h es m a l l e f f e c t ss e e ni nt h es t a t i cw i n d - t u n n e lt e s t sd i s c u s s e de a r l i e r .

The f l i g h tc h a r a c t e r i s t i c so b s e r v e d when t h e 40° nosestrakes were added a l s o correlatedwellwiththestaticwind-tunneldata, which showed t h a tt h es t r a k e s produced a largedegradation i n d i r e c t i o n a l s t a b i l i t y above l o o angle of a t t a c k .

(See f i g . 21 .) I nt h ef r e e - f l i g h tt e s t s ,t h i se f f e c t was manifested i n the form of "nosewandering" i n yaw which became q u i t es e v e r e above about 20° angle of a t t a c k .

The model couldnot be flownabove 25O angle of attackbecause of t h i s problem.

Thus,the 40° nose s t r a k e s do notappearto be a n a c c e p t a b l es o l u t i o nt ot h e wing- rockproblem.

The f r e e - f l i g h t results f o rt h ee f f e c t s of c.g. l o c a t i o n and s t r a k e - f l a ps e t t i n g canbest be analyzed by examiningthe s t a t i c trimmed-canard v a r i a t i o n sf o rt h e s e cases.Presented i n f i g u r e 43 a r e static trimmed-canard variationswithangle of a t t a c kf o rt h eb a s e l i n ec o n f i g u r a t i o n (6, = 30° and c.g. = -0.12c) and €or two othercombinations of s t r a k e - f l a ps e t t i n g and c.g.location. The d a t a show s i g n i f i - c a n tc h a n g e si ns t a t i c trimmed-canard settingsbecause of a 5-percent a f t c.g. s h i f t o r a 30° change i ns t r a k e - f l a ps e t t i n g . As d i s c u s s e de a r l i e r ,c a n a r ds e t t i n g had a s u b s t a n t i a l e f f e c t on s t a t i c d i r e c t i o n a l s t a b i l i t y and ruddereffectiveness.

Unloading the canard had t h e f o l l o w i n g e f f e c t s : ( 1 ) It degraded C be low "B a = 30° and enhanced above 30°; and ( 2 ) it degraded C throughoutthe "6r test angle-of-attackrange. The n e t e f f e c t on d i r e c t i o n a l s t a b i l i t y is shown i n f i g - ure 44, which compares trim f o rt h et h r e ec o n f i g u r a t i o n sd i s c u s s e dp r e v i o u s l y .

c"B These same e f f e c t s were o b s e r v e di nt h ef r e e - f l i g h t tests of t h e a f t c.g.location (-0.07c) and/orofzerostrake-flapdeflection. For example, fortheworst-case configuration combining aft c.g. and 6 s = O o ( c o n f i g u r a t i o n 31, nose wandering

i n yaw (degradationof C ) w a s observedtobegin a t a lower angle of a t t a c k and t o

"B be more severe below a = 30° t h a nt h a tf o rt h eb a s e l i n ec o n f i g u r a t i o n . However, above a = 30° a reduction in nose wandering was observed, which r e f l e c t e dt h e h i g h e rl e v e l s of as compared with those of thebaselineconfiguration.(See c"B f i g . 44.) Nonetheless, the degraded rudder effectiveness of c o n f i g u r a t i o n 3 gener- a l l yl e dt o loss of c o n t r o la ta n g l e s of a t t a c k somewhat lowerthan 4 0 ° .

S U M M A R Y OF RESULTS The results of a w i n d - t u n n e lf r e e - f l i g h ti n v e s t i g a t i o n of thehigh-angle-of- a t t a c k c h a r a c t e r i s t i c s of a forward-swept-wing f i g h t e r - a i r p l a n ec o n f i g u r a t i o n may be summarized asfollows: 1 . The model e x h i b i t e d good dynamic s t a b i l i t y c h a r a c t e r i s t i c s up t o 40° angle of a t t a c kw i t hs t a b i l i t ya u g m e n t a t i o n i n p i t c h ,r o l l , and yaw. F l i g h ta th i g h e r angles of a t t a c k was notpossiblebecause of lack of yaw-controleffectiveness.

2. Because of theveryhighlevel of s t a t i c p i t c h i n s t a b i l i t y of theairframe, c o n t r o l l e d f l i g h t was notpossiblewithoutthepitchstabilityaugmentationsystem ( S A S ) . With the SAS a c t i v a t e d ,t h e model e x h i b i t e de x c e l l e n tl o n g i t u d i n a lf l y i n g c h a r a c t e r i t i c s where adequatepitchcontrol was a v a i l a b l e . These r e s u l t si n d i c a t e t h a t it i s feasibletodesignforveryhighlevels of a i r f r a m ep i t c hi n s t a b i l i t y a s longasthere is adequatecontrol power and sensorinformationtoprovidethe necessarylevel of s t a b i l i t ya u g m e n t a t i o n .

3 . Nose-up p i t c hc o n t r o l below about 20° angle of a t t a c k was found t o be marginal because of c a n a r ds t a l l . Making t h es t r a k ef l a p a n a c t i v ep i t c h - c o n t r o l d e v i c es h o u l da l l e v i a t et h i sc o n d i t i o n .

4 . With t h er o l ls t a b i l i t ya u g m e n t a t i o n s y s t e m d e a c t i v a t e d ,t h e model e x h i b i t e d large-amplitude undamped r o l lo s c i l l a t i o n s (wing rock) above 25O angle of a t t a c k which made controlmarginal. The r o l l SAS actingthroughthepowerfulflaperons e f f e c t i v e l ys u p p r e s s e dt h e wing rock up t ot h eh i g h e s ta n g l e of attackachieved (40'1.

5. The s t a t i cl a t e r a l - d i r e c t i o n a ls t a b i l i t yc h a r a c t e r i s t i c s of theconfigura- t i o n were notdominated by any one configurationfeaturethroughouttheangle-of- attackrange. Above about 30° angle of attack,thefuselageforebodyproduced b e n e f i c i a l c o n t r i b u t i o n s t o s t a t i c d i r e c t i o n a l s t a b i l i t y , b u t it alsoproduced unstablevalues of damping i n r o l l . The forward-swept wing was primarilyresponsible f o r t h e l a t e r a l s t a b i l i t y c h a r a c t e r i s t i c s up t oa b o u t 30° angle of a t t a c k ; however, t h ec h a r a c t e r i s t i c s above 30° angle of attackappearedto be s t r o n g l yi n f l u e n c e d by an i n t e r a c t i o n between theforebody and thecanard-wingflowfields.

6. Canard s e t t i n g was found t o i n f l u e n c e s i g n i f i c a n t l y the high-angle-of-attack, vertical-tail f l o w c h a r a c t e r i s t i c sa n dh e n c e static d i r e c t i o n a l s t a b i l i t y , yaw damping,andruddereffectiveness.

LangleyResearchCenter NationalAeronauticsandSpaceAdministration Hampton, VA 23665 December 2, 1983 REFERENCES 1. Krone, Norris J., Jr.: Divergence Elimination With Advanced Composites. A I A A Paper N o . 75-1009, Aug. 1975.

2. Nguyen, Luat T.; Yip, Long P.; and Chambers, Joseph R.: Self-Induced Wing Rock of Slender Delta Wings. AIAA-81-1883, Aug. 1981.

3. Chambers, Joseph R.; Bowman, James S., Jr.; and Malcolm, Gerald N.: S t a l l / S p i n Test Techniques Used by NASA. S t a l l / S p i n Problems of M i l i t a r yA i r c r a f t , AGARD-CP-199, June 1976, pp. 13-1 - 13-12.

4. Grafton, Sue B.; Chambers, Joseph R.; and Coe, Paul L., Jr.: Wind-Tunnel Free- F l i g h t I n v e s t i g a t i o n of a Model of a Spin-ResistantFighterConfiguration.

N A S A TN D-7716, 1974.

5. Bates, William R.: Static S t a b i l i t y of Fuselages Having a R e l a t i v e l y F l a t Cross Section. NACA TN 3429, 1955. (Supersedes NACA RM L9106a.1 6. Chalk, C. R.; Neal, T. P.; Harris, T. M.; P r i t c h a r d F. E.; and Woodcock, R. J.: Background Informationand User Guide for MIL-F-8785B(ASG), " M i l i t a r y S p e c i f i c a t i o n - F l y i n gQ u a l i t i e s of PilotedAirplanes." AFFDL-TR-69-72, U.S. A i r Force, Aug. 1969. (Available from DTIC as AD 860 856.1 7. Etkin, Bernard: Dynamics of Atmospheric Flight. John Wiley & Sons, Inc. , c. 1972.

8. Nguyen, Luat T.; G i l b e r t , William P.; and Ogburn, Marilyn E.: Control-System Techniques for Improved Departure/SpinResistance for F i g h t e r Aircraft. N A S A TP-1689, 1980.

9. Moul, Martin T. ; andPaulson,John W. : Dynamic Lateral Behaviorof High- NACA RM L58E16, 1958.

Performance Aircraft.

TABLE I .. MODEL MASS AND GEOMETRIC CHARACTERISTICS

Weight. lb ................................................................... 61.75

Moments of i n e r t i a : Ix. s l u g - f t 0.484 ...............................................................

Iy. s l u g - f t 4.49 ...............................................................

Iz. s l u g - f t 4.61 ...............................................................

Wing:

Span. f t ................................................................... 4.35

Mean aerodynamic chord. f t ................................................. 1.15

Area. f t 2 .................................................................. 4.74

Aspect r a t i o ............................................................... 4.0

Sweep ( l e a d i n ge d g e ) .d e g .................................................. -29.27 Sweep ( 0 . 2 5 ~c h o r d ) .d e g ................................................... -33.73

Taper r a t i o ................................................................ 0.404

Flaperon area ( t o t a l ) . f t 2 ................................................. 0.734 F l a p e r o nd e f l e c t i o n .d e g .......................... 17.5 (T.E.D.) t o -27.7 (T.E.U.)

Canard:

Span. f t ................................................................... 2.18

Area ( t o t a l ) . f t 2 .......................................................... 0.947

Aspect r a t i o ............................................................... 1.47

T a p e r r a t i o ................................................................ 0.318

D e f l e c t i o n . d e g ....................................... 30 (T.E.D.) t o -60 (T.E.U.)

Vertical t a i l :

Height. f t ................................................................. 1.07

Area. f t 2 .................................................................. 0.864

A s p e c t r a t i o ............................................................... 1.32

Taper ratio ................................................................ 0.306

Rudder area. f t 2 ........................................................... 0.177

R u d d e rd e f l e c t i o n .

deg ..................................................... *30

S t r a k e f l a p :

Area ( t o t a l ) . ft2 .......................................................... 0.267

D e f l e c t i o n . d e g ............................................................ *30 TABLE 11.- SUMMARY O F FREE-TO-ROLL RESULTS Angle of a t t a c k , C o n f i g u r a t i o n R e s u l t s de9 Completeconfiguration, 20 L i g h t l y dampeda b 6 = -25O Mildwingrock C ModeratewingrockC d 35 Large-amplitude wing rock d Large-amplitudewingrock d 45 Large-ampli tude wing rock d Body, wing, and t a i l ( c a n a r do f f ) Large-amplitudewingrock d 4 0 Body andwing (canardand tail off ) Large-ampli tude wing rock d Body and t a i l (canardandwingoff) 4 0 Large-amplitudewingrock b Mildwingrock Completeconfiguration,plus 40 40° n o s e s t r a k e s ; 6, = -25O ~ " - " a L i g h t l y damped: R o l lm o t i o n se v e n t u a l l yd i e Out.

'Mild wing rock: A $ < fl5O.

'Moderate wing rock: f l < A $ < f15' dLarge-amplitudewingrock: A$ > +25 O .

. . . . " . . .. . ._ - .. " .... . . . .

Figure 1.- System of axes.

I 4.353

+

.593

( a ) Plan-viewsketch of f l a t - p l a t e wing model.

Figure 2 .- Drawings of models.Dimensions aregivenin feet unlessotherwisespecified.

NOSE STRAKE

e

CANARD

i

STRAKE

L%"

4 7.68 ft +

1 - 4.35ft - 4

RUDDER fl

( b ) Three-view sketch of b a s i c model.

F i g u r e 2.- Concluded.

N W E83-127 (a) Side view.

Figure 3.- Photographs of basic model.

L83-128 (b) Mounted i n t h e Langley 30- by 60-FootTunnelduringstatic-forcetesting.

Figure 3.- Concluded.

L l

7 "

( a ) Rear location.

(b) Side location.

( c ) Top location.

Figure 4.- Locations of spin-chutecanister.

. 5

A-A

Figure 5.- Geometry of 40° nosestrakes. Dimensions a r eg i v e ni n i n c h e s u n l e s s o t h e r w i s e s p e c i f i e d .

Wing reference line

Position B, deg

c, deg

12.3

Maximum T.E. D e 13.6 17.5

0 0 0

U ndef I ect ed

- 19.9

-19.3 -27.7 Maximum T.E. U.

Figure 6.- Flaperon-deflectiondefinition. All a n g l e s are measuredperpendicular to 0 . 7 5 ~ .

Figure 7 . - Photograph of basic model i n free f l i g h t .

h) W W Figure 8.- Test s e t u p for wind-tunnel free-flight tests.

.q . 2 cm - . 2 - .q - -6 0 1 0 0 30 A Off -10 0 10 20 30 90 50 60 70 80 90 100 a, deg (a)Positive canarddeflections.

Figure 10.- Effect of canarddeflection on static longitudinal characteristics.

6 = 17.5O; 6 = 30°.

f S -.6 - .8 2.8 2.6 2.0 1.8 c~ 1.6 and l . q ' D 1.2 1 .o .8 .6 .2 (b) Negative c a n a r dd e f l e c t i o n s .

Figure 10. - Concluded.

-10

deg

c, t’

-20

-30

-40

- 50

-60

0 30 40 50 60 . 70

a, deg

Figure 1 1 . - Canard deflection required €or trim. c . g . a t -0.07;; 6, = 17.5O; 6 , = 30°.

.6 A .2 - . 2 -.S - . 6 - .8 bS# deg 0 30 0 0 2.8 0 -30 2 -6 2 .S 2.2 2 .o 1.8 C 1.6 L and 1 .q c D 1.2 1 .o .8 . 6 . 9 .2

0 T I I I I I I I I I I I I I I I I I I I ~

-10 0 10 20 30 SO 50 60 70 80 90 100 a, deg Figure 12.- Effect of strake-flapdeflection on s t a t i c longitudinal characteristics. 6 c = -25 O ; 6f = 17.5 O .

2.6 2.11 2.2 2.0 1.8 1.6 cL

1 .'4

a n d 1.2 c D 1 .o .8 .6

.'4

.2 -10 0 10 20 30 '40 50 60 70 80 30 100 a, deg Figure 13.- Effect of flaperondeflection on static longitudinal characteristics. 6, = -25O; 6 , = 30°.

.04

C

ys 0

- .04

* 008 .006 -004

c"

f3 .002 - ,002 - ,004 Vertical ta I i I 0 On 0 Off ( a ) 6, = Oo.

Figure 14.- Effect of v e r t i c a l t a i l on s t a t i c 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 .

6, = 17.5O; 6 , = 30°.

.011

C

yP 0

- ,011

.010 ,008 .006

c .ooq

“0

Vertical tail .002 0 On Off - ,002

- . 0011

. ooq

.002

cb

- ,002

- . ooq

- - 006

-10 0 10 20 30 q0 50 60 70 80 90 100 a, deg ( b ) 6, = -25’.

Figure 14. - Concluded.

0 X-29A; 6, = -25O ; bf = 17.5O ; 6 = 30°

S

n Flat-plateforward-sweptwing

- ooq

“006

Figure 15.- Comparison of lateral stability characteristics of X-29A and f lat-plate f orward-swept wing.

. 0 1 1

C

yB 0

- .0q

.010 -008 .006

. OOLI

C

" P

.002 - ,002

- . OOLI

.002 - .002

%

- .ooq

Figure 16.- Effect of canarddeflection on s t a t i c l a t e r a l i l i r e c t i o n a l s t a b i l i t y .

6, = 17.5O; 6, = 30°.

I

( a ) Canard o f f .

(b) Canard on. 6, = O O .

6 8 3 - 1 31 Figure 17.- Tuft-flowvisualization of X-29A v e r t i c a l t a i l a t a = 25O.

6 , = 17.5O; 6 , = 30°.

.0q

C

yB 0

- .0q

.008 .006 .ooq

C

"P

.002 .002 .002

czB -.002

- .ooq -10 0 10 20 30 q0 50 60 70 80 90 100 a, deg .010 -008 .006

C

. OOY

"B

.002 .002 ,002

- -002

- .ooq

- - 006

-10 0 10 20 30 q0 50 60 70 80 90 100 a, deg F i g u r e 19.- E f f e c t of s t r a k e - f l a pd e f l e c t i o n on s t a t i c 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 . 6c = -25O; 6f = 17.5O.

. 0q

C

ye 0

- .0q

-008 .ooq

Canister

C

0 Off .002

Rear

0 Side

A Top - .002 .002

Cb -.002

- .ooq -10 0 10 20 30 90 50 60 70 80 90 100 Figure 20.- Effect of spin-chutecanisterlocation on staticlateral-directional s t a b i l i t y . 6c = -25"; 6 , = 17.5"; 6 s = 30".

C

yP

- .Orl

ooq

alp sensor

0 Off

C . 002

0 O n

np

- .002

- ooq

- 002

- .ooq

- =006

-10 0 10 20 30 YO 50

a, deg

.Figure 22.- Effect of boom-mounted a/@ vanesensoron s t a t i c 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 . 6, = -25O; 6f = 17.5O; 6 , = 30°.

Figure 23.- Helium-bubble flow v i s u a l i z a t i o n of X-29A forebody flow a t h i g h a n g l e s of attack.

.05

AcY 0

.02 .01 - .01 .07 .06 .05

. OLf

.03 .02 .01 - .01 -10 0 10 20 30 t10 50 60 70 80 90 100

a , deg

Figure 24.- Effect of ailerondeflection. 6, = -25O; 6, = 3 0 ' .

.05 - -05

-. 10

.02 .01 -.01 - .02 - .03 - .oq .08 .07 .06 .05 Acl .oq .03 .02 .01 "40 50 60 70 80 30 100 -10 0 10 20 30 a, deg Figure 25.- Effect of canard d e f l e c t i o n on a i l e r o ne f f e c t i v e n e s s . 6 = -22.6O a ( b f , l e f t = 17.5O; 6 f , r i g h t = -27.7O); 6 , = 30°.

.05 AC -.OS -.lo "15 brJ d~ 0 -30 0 -15 .06 .os .09

- 03

.02 .01 - .01 - .02 .01 Acl - .01 - .02 - . l o 0 10 20 30 q0 50 60 70 80 90 100 a, deg Figure 26.- Effect of rudder d e f l e c t i o n . 6, = -25O; d f = 17.5O; 6 , = 30°.

.05 AC -.OS -.lo "15 .07 .06 .OS

. OLi

.03 hCn .02 .01 -.01 - .02 .01 Acl -.01 - .02

~ I I I I I I I I I I I I I I I I I I I

- -03 - 1o 20 30 L ~ O 50 60 70 80 90 loo a, deg Figure 27.- Effect of c a n a r dd e f l e c t i o no nr u d d e re f f e c t i v e n e s s . 6, = - 3 O O ; 6 , = 17.5O; 6 , = 3 0 ° .

6 , I deg

0 -10 0 -25

0 -50

D Off -Y -8 Unstable

* Stable

-10 -20 -30 -10 0 10 20 30 40 50 60 70 80 90 100 a, deg F i g u r e 28.- E f f e c to fc a n a r dd e f l e c t i o n on d y n a m i cl o n g i t u d i n a ls t a b i l i t y d e r i v a t i v e so b t a i n e dd u r i n gp i t c h i n g - o s c i l l a t i o n tests. 6f = 200; 6 s = 30°; A9 = +50.

C + C sina

n

P ni

.8 Unstable

Cl + C . sina

P ZB

3 Stable

- .q -10 0 10 20 30 q0 50 60 70 a, deg F i g u r e 29.- E f f e c t of c a n a r dd e f l e c t i o n on dynamic 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 d e r i v a t i v e so b t a i n e dd u r i n gr o l l i n g - o s c i l l a t i o n tests. 6, = 2 0 ° ; 6 s = 30°; A + = *IO0.

0- X-29A; 6 = -25' ; bf = 20' ; b = 30'

C S

Flat-plate forward-swept wing

.2

+

Cz- sin a

c z P P

- -2

- * 6

-1 0 0 10 20 30 YO 50 60 70

a, deg

Figure 30.- Comparison of roll-dampingcharacteristics of X - 2 9 A and f l a t - p l a t e forward-swept wing. A $ = & I O 0 .

c y P + c% sin a 0

- - q I L L L L l I I I I I I T T I I I I

40' nose strake 1.2

0 O n

0 Off

.8 C I sin a

Cnp +

n b .Ll

- . q . q Unstable Stable - .Y -.8 -10 0 10 20 30 q0 50 60 70 a, deg Figure 31 .- E f f e c t of 40° nosestrakes on dynamic 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 d e r i v a t i v e so b t a i n e dd u r i n gr o l l i n g - o s c i l l a t i o n tests. 6= = -25O; 6f = 2 0 ° ; g S = 30°; A $ = &So.

II

-. c COS a Vertical tail

yf3 0 On 0 Off Unstable Stable F i g u r e 32.- E f f e c t of v e r t i c a l tail on dynamic 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 d e r i v a t i v e s o b t a i n e d d u r i n g y a w i n g - o s c i l l a t i o n tests. 6c = -25O; 6f = 2 0 ° ; 6 s = 30°; AJI = * S o .

i j C 3 deg

Cy, - C - cos a

0 -10

yP

0 -25

O -50

A Off . L f Unstable Stable

- . L t

-1.2 -10 0 10 20 30 LfO 50 60 70 a, deg Figure 33.- E f f e c t of canarddeflection on dynamic 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 derivatives obtained during yawing-oscillation tests. 6, = 2 0 ° ; 6 s = 30°; AJI = f5O.

1 . 5 Des i red

~ ” -

1.0 \ \ j w , \

rad/sec 3

\ . 5 \ \ \ V I .

CJ, rad/sec ( a ) Unaugmented. t2 = 0.85 sec.

1.5 1.0 j w , rad/sec . 5

-1.5 -1 .o - . 5 0 . 5 1 .o

u , rad/sec ( b ) Augmented. wn = 1.17 rad/sec; < = 0.65; tl,2 = 0.91 sec.

Figure 34.- Root-locus representation of pitch-stability-augmentation, short-periodcharacteristicswith a : = 3 0 ° . Values shown are f o r full-scale airplane.

j w , rad/sec

-1.5 - 1 .o -.5 0 .5 1 .o

u, rad/sec (a) Unaugmented. t2 = 1.20 sec.

I I

-1.5 -1 .o -.5 0 .5 1 .o

u, rad/sec = 1.65 sec.

(b) Augmented. wn = 1.63 rad/sec; 5 = 0.26; t f /2 Figure 35.- Root-locus representation of roll-stability-augmentation, Dutch roll characteristics with a = 30°. Values s h a m are €or full-scale airplane.

6 8 3 - 1 3 3 Figure 36.- S t a b i l i t ya u g m e n t a t i o ns y s t e m .

Basic

ARI on (KAR, = 2.3)

- 06

ICDP .03

- a 0 1

0 5 10 15 20 25 30 35 LfO Li5 50

a, deg

Figure 37.- Effect of A R I on lateralcontroldivergenceparameter. 6, = Trim; 6 , = 17.5O; 6 , = 30°.

= OLlO

= 035

.030

- 025

C

n

p, dY n .020

-015

-010

0 10 20 30 LlO 50 60 70

a, deg

w i t ha n g l e of a t t a c k . 6c = Trim; F i g u r e 38.- V a r i a t i o n of C np,dyn 6 , = 17.5O; 6, = 3 0 ° .

I

2 . 0

use of

1.5

c.L

1.0

.5

Figure 39.- Model lateral-directional flight characteristics with full augmentation.

Q, deg 20

t

0 1 I I J

-

-

F u l l u p

n A /

L I I L - t L I

Full down

0 5 10 15 0 5 10 15 t, sec t, sec ( a ) Baseline SAS. (b 1 Bduced-gain SAS .

Figure 40.- Effect of p i t c h SAS on model f l i g h t c h a r a c t e r i s t i c s . Values shown are i n model scale.

2.0

Control marginal because of

1.5

large-amplitude wing rock

cL

- Undamped roll oscillations

increase pilot work load

1.0

- Model easy to fly;

roll motions lightly damped

. 5

0 10 20 30 40 50

a, deg

Figure 41.- Model lateral-directionalflightcharacteristicswithout r o l l augmentation.

10 -

0 . 1 I I I I I 1

I

r

I I I * I 1

-200 1

I

I

80 -

I

40 - I

r y deg/sec A n

A I\ - A -

-

-

W

-

-40 I I I I I -80

20 -

10 -

0 -

-10 -

- 20

0 5 10 15 20 25 t, sec Figure 42.- Effect of r o l l SAS on model flight characteristics. Values shown a r e i n model scale.

6 , deg cog. location Configuration

S

30 -0.07C 1

""

30 -0.12c 2 (baseline)

"-

0 -0.07e 3

\

- 10

6 deg

c, t'

\ \

- 20

- 30

- 40

- 50

\ I I I I I 1

- 60

0 10 20 30 40 50 60

a, deg

Figure 43.- Effect of c.g. location and strake-flapdeflection on 6c,t; 6 , = 17.5O.

6 deg c.g. location Configuration

S

0 30

-0.07t

0 30 2 (baseline)

-0.12t

0 0

-0.07C

- 002

0 10 30 LlO 50 60 70

a, deg

Figure 44.- E f f e c t of s t r a k e - f l a pd e f l e c t i o n and c.g. location on d i r e c t i o n a l s t a b i l i t y . 6, = T r i m ; 6, = 17.5O.

~~ 1. Report No. vernment Accession No. 3. Recipient's Catalog No.

NASA TP-2230 4. Title and Subtitle 5. Report Date WIND-TUNNEL FREE-FLIGHT INVESTIGATION

I F e b r u a r y 1984

OF A MODEL OF A FORWARD-SWEPT-WING FIGHTER 6. Performing Organization Code CONFIGURATION

I 505-43-1 3-01

7. Author(s) 8. Performing Organization Report No.

D a n i e l G. Murri, Luat T. Nguyen, L-15602 and Sue B. G r a f t o n 9. Performing Organization Name and Address NASA LangleyResearchCenter 11. Contract or Grant No.

Hampton, VA 23665

I

13.Type of Report and Period Covered T e c h n i c a lP a p e r 12. Sponsoring Agency Name and Address NationalAeronauticsandSpaceAdministration 14. Sponsoring Agency Code Washington, DC 20546

r"--

15. Supplementary Notes 16. Abstract A w i n d - t u n n e lf r e e - f l i g h ti n v e s t i g a t i o n w a s conducted t o s t u d yt h ed y n a m i cs t a b i l i t y c h a r a c t e r i s t i c s of a model of a f o r w a r d - s w e p t - w i n gf i g h t e r - a i r p l a n ec o n f i g u r a t i o n a t h i g ha n g l e s of a t t a c k .V a r i o u so t h e rw i n d - t u n n e lt e c h n i q u e se m p l o y e di n the s t u d y i n c l u d e d static- anddynamic-(forced-oscillation) force tests, f r e e - t o - r o l l tests, a n df l o w - v i s u a l i z a t i o n tests. A u n i q u ef a c e t of t h es t u d y w a s t h ee x t r e m el e v e lo f s t a t i c p i t c h i n s t a b i l i t y ( i n e x c e s s of negative32-percent s t a t i c m a r g i n )i n h e r e n ti n t h ea i r f r a m ed e s i g nw h i c hp r e c l u d e df r e e - f l i g h tt e s t i n gw i t h o u ts t a b i l i t ya u g m e n t a - t i o n i np i t c h .R e s u l t s are presentedwhichemphasizethehigh-angle-of-attack aero- dynamics and the vehicle-component contributions t o t h e s ec h a r a c t e r i s t i c s . The e f f e c t s o ft h e s ea e r o d y n a m i cc h a r a c t e r i s t i c s on t h eh i g h - a n g l e - o f - a t t a c kf l y i n g q u a l i t i e s o ft h ec o n f i g u r a t i o n are d i s c u s s e di n terms of r e s u l t s ofthewind-tunnel f r e e - f l i g h t tests.

7. Key Words (Suggested by Author(s1) 18. Distribution Statement

- Unlimited

Forward-sweptwing U n c l a s s i f i e d Relaxed s t a t i c s t a b i l i t y Dynamic s t a b i l i t y A i r c r a f t s t a b i l i t y a n dc o n t r o l S u b j e c tC a t e g o r y 08 High a n g l e of a t t a c k 9. Security Classif. (of this report] 20. Security Classif. (of this page) 21. No. of Pager 22. Price U n c l a s s i f i e d i f i e d A 0 4 For sale by the National Technical Information Service, Springfield, Virginla 22161 NASA-Langley, 1984 Nationa.1 Aeronautics and THIRD-CLASS BULK RATE Postage and Fees Paid National Aeronautics and Space Administration Space Administration NASA451 Washington, D.C.

" USMAIL 20546 Official Business Penalty for Private Use, $300 POSTMASTER: If Undeliverable (Section 1 5 8 Postal Manual) Do Not Return, J