What was the main focus of the wind-tunnel investigation?

The main focus was to study the stability and control characteristics of a forward-swept-wing fighter configuration at high angles of attack.

What challenges were encountered during free-flight testing?

The extreme level of static pitch instability required stability augmentation, which precluded free-flight testing without it.

How did the model perform in terms of dynamic stability?

The model exhibited good dynamic stability characteristics up to a 40° angle of attack, but testing could not be conducted beyond this due to lack of yaw control effectiveness.

19840009116 — Wind-tunnel free-flight investigation of a model of a forward-swept-wing fighter configuration

Q: What factors influenced lateral-directional stability?

Lateral-directional stability was primarily influenced by the forward-swept wing and vertical tail at low to moderate angles of attack, with significant component interactions also affecting the characteristics.

Q: What was the role of the stability augmentation system?

The stability augmentation system was effective in providing the required level of stability as long as there was sufficient pitch control available.

Document

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NASA

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Technical

Paper

February 1984

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Wind-Tunnel Free-Flight

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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

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 = -

' n a s

6a a " ="- - a6 ' n

6 r r 6 r a 6r

acY

a ' n

c z = - cn = - c y = -

Pb Pb Pb

P 3 - P a -

P a -

2v 2v 2v

a ' n acY

= -

CI, = - cy. = -

'n; !b

B ib B bb

a -

a - a -

2v 2v 2v X, acY

c1 = - cy = -

rb rb

r a -

r 3 - 2v 2v

a ‘ m acN

cm = - CN = - CA = -

qc qc 9E

g a - q a - g a -

2v 2v 2v

a ‘ m

c = -

m - .-

a ac

a -

2v IZ = C c o s a - - C , s i n a ‘ n

P , dYn IX B

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

CANARD

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

ys 0

- .04

* 008 .006 -004

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

yP 0

- ,011

.010 ,008 .006

c .ooq

“0

Vertical tail .002 0 On Off - ,002

- . 0011

. ooq

.002

- ,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°

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

yB 0

- .0q

.010 -008 .006

. OOLI

" 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°.

( 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

yB 0

- .0q

.008 .006 .ooq

.002 .002 .002

czB -.002

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

. OOY

.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

ye 0

- .0q

-008 .ooq

Canister

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".

- .Orl

ooq

alp sensor

0 Off

C . 002

0 O n

- .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

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

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.

-. 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

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

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 ° .

2 . 0

use of

1.5

c.L

1.0

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

Q, deg 20

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

- 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 I I * I 1

-200 1

80 -

40 - I

r y deg/sec A n

A I\ - A -

-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

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

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

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

19840009116 · Wind-tunnel free-flight investigation of a model of a forward-swept-wing fighter configuration

Overview

Key points

Document

NASA

Technical

Paper

February 1984

Wind-Tunnel Free-Flight

Technical

Paper

Wind-Tunnel Free-Flight

Investigation of a Model

Fighter Configuration

Langley Research Center

Hampton, Virginia

NOSE STRAKE

CANARD

STRAKE

Wing reference line

Position B, deg

Maximum T.E. D e 13.6 17.5

U ndef I ect ed

Canister

Rear

0 Side

* Stable

C + C sina

3 Stable

Flat-plate forward-swept wing

Basic

ICDP .03

= OLlO

Full down

- Undamped roll oscillations

- Unlimited

Source & rights

Document details