Document
NASA MEMO 10-7-58L c
MEMORANDUM
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FLIGHT MEASUREMENTS O F SOME O F THE FLYING QUAUTLES AND STABILITY DERIVATIVES O F A SUPERSONIC FIGHTER AIRPLANE By Christopher C. Kraft, Jr., Milton D. McLaughlin, Jack A. White, and Robert A. Champine Langley Research Center Langley Field, Va.
c Y ’ v I NATIONAL BONAUTICS AND SPACE ADMINISTRATION I 1 .
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FLIGHT MEAStIREMENTS OF SOME OF THEFLYING QUALITIES AND STABILITY DERIVATIVES OF A SUPERSONIC FIG€E'EB AIRPLANE* By Christopher C. Kraft, Jr., Milton D. McLaughlin, Jack A . White, and Robert A . Champine SUMMARY Flight measurements have been made to determine the flying qualities and some of the stability derivatives of a supersonic fighter airplane.
The results are presented in the form of measured flight data and pil opinion.
The damping of the short-period longitudinal oscillations is some- what low. The feel forces provided by the longitudinal feel system are considered good by the pilots. The longitudinal-control-system characteristics that result from the nonlinear gearing between the stick and stabilizer result in poor handling characteristics for all indicated airspeeds. It appears as though a more linear stick-to-stabilizer rela- tionship near trim would result in improved flying qualities throughout the flight regime. Also, the longitudinal stick-fixed stability as measured by the variation of stick position with normal acceleration is adversely affected by structural deformation during accelerated maneu- vers. The airplane has a high roll-to-yaw ratio but one which is within the present flying-qualities requirements. The pilots dislike the longi- tudinal trim system because of the difficulty experienced when trying trim precisely and the overshoot which occurs when making a large or rapid trim correction. The roll performance of the airplane is con- sidered adequate for Mach numbers below 0.9, but the performance dete- riorates rapidly in the high Mach number, high-dynamic-pressure region.
INTRODLJCTION This paper presents an investigation of the flying qualities and measurements of some of the stability derivatives of a supersonic day fighter for both carrier-based and land-based operations. Flight tests , %itle, Unclassified.
" b+ to measure the flying qualities and other characteristics of the air- plane are presented in references 1 to 10. Tests of modern airplanes such as the one in the present investigation are needed to extend the F .
present flying-qualities specifications of reference 11 to the new flight regimes covered by the high performance capabilities of this type of aircraft.. In addition, the need for flying-qualities inves- tigations is continuous to ascertain if there is a need for additional requirements to or revisions in the present requirements. This fighter airplane was extensively tested during the design stage by.both wind- tunnel and rocket-model techniques. (For example, see refs. 12 to 16. ) It is therefore interesting and beneficial to future design to contin the tests of this particular airplane in flight so that previous tests can be compared with flight test results.
The test airplane incorporates several new design features in it external geometry which make the airplane of general interest. Such features are leading-edge chord-extensions, leading-edge droop, high wing and low tail, and a variable-incidence wing to improve take-off and landing characteristics. A l s o , the longitudinal control system of this airplane combines such features as a spring, stick dampers, bob- weights sensitive to both normal and pitching accelerations to provide force feel to the pilot, and a nonlinear linkage combined with an irre- versible power control system.
This report deals with the first phase of the flight investigat of the test airplane and discusses some of the handling qualities of airplane that were obtained during pilot evaluation flights. A l s o , some brief test maneuvers have been made to determine some of the airplan stability derivatives.
SYMBOLS r lateral acceleration b wing span
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C mean aerodynamic chord of wing Rolling moment rolling-moment coefficient, C l qSb normal-force coefficient, CN Pitching moment pitching-moment coefficient,
qse
‘ m Cn
CY
c y c l e s t o damp t o one-halfamplitude c1/2 acceleration due t o gravity pressure altitude moment of i n e r t i a of airplane about X s t a b i l i t y a x i s i moment of i n e r t i a of airplane about Y s t a b i l i t y a x i s moment of i n e r t i a of airplane about Z s t a b i l i t y a x i s product of i n e r t i a r e f e r r e d to X and Z s t a b i l i t y axes Mach number mass of airplane normal acceleration, g u n i t s Pb - helix angle 2v r o l l i n g v e l o c i t y P P period of o s c i l l a t i o n dynamic pressure or nondimensional pitching velocity wing area time t o damp t o one-halfamplitude true airspeed equivalentairspeed airplane weight CLV vane angle P sideslip angle 6a aileron deflect ion
4- rudder deflection
B bank angle
rolling parameter, - 57.3 e!
ve P A increment Stability derivatives are indicated by subscript notation; for example, Rotary derivatives are defined as indicated by the following: Subscripts: i d C calibrated f fuselage W wing Dot over quantity indicates differentiation with respect to time.
DESCRIPTION OF AIRPLANE The Chance Vought F8U-1 airplane is a high-wing, l o w - t a i l f i g h t e r airplane intended for both carrier- and land-based operations and designedforuse as a supersonicdayfighter. The airplane powerplant i s a P r a t t & Whitney 557-P-4 with afterburner. Pictures.of the test airplane are shown i n f i g u r e 1, a drawingof t h e airplane is g i v e n i n f i g u r e 2, and pertinent characteristics of the airplane are presented i n t a b l e I.
The t e s t a i r p l a n e has a variable-incidence wing for use during landingandtake-off. The wing is moved hydraulically to an incidence of 7' in the landing condition, and in the clean condition the wing is fixed a t -lo. The wing is equippedwith a leading-edge f l a p ( c a l l e d leading-edgedroop)whichcan a l s o be operated hydraulically to three d i f f e r e n tp o s i t i o n s . These positions are clean,cruisedroop, and landingdroop. The droop on eachsideofthe wing is composed of two sections, one sectionextending from the root to the leading-edge chord- extension (inboard section) and the other section extending from this p o i n tt ot h e wing tip(outboardsection).(Seefig. 2 . ) When t h e wing i s raised to the landing position, the inboard leading edge is drooped 25O and theoutboardleading edge i s drooped 27O. Also, it is possible t o p u t t h e droop in the landing position with an emergency air system.
A f t e r t h i s i s done, t h e droop s t a y s i n t h e l a n d i n g p o s i t i o n r e g a r d l e s s of wing position.Inadditiontotheleading-edge droop forlanding, t h e a i l e r o n s a r e d e f l e c t e d down 20' as a f l a p and t h e small f l a p s at t h e wing rootaredeflected down 20°. When t h e wing i s raised,thehorizon- t a l t a i l i s automatically deflected 5 O leading edge up t o minimize t h e changes i n trim. The leading-edgedroopcan be deflected 6 . 8 O and 70 (inboard and outboardsections,respectively)intothecruise-droop p o s i t i o n t o improve cruise and maneuver performance a t subsonic and transonicspeeds.
The control surfaces of the airplane are a l l hydraulicallyoperated with irreversible systems and t h e f e e l f o r c e s t o t h e p i l o t a r e s u p p l i e d by a r t i f i c i a l means. The aileron- and rudder-control f e e l forces are suppliedbysimplesprings. The forcesrequired and thedeflection ranges available in the aileron and ruddercontrolsystemsintheclean and landingconditionsaredifferent. The c h a r a c t e r i s t i c so ft h ea i l e r o n and ruddersystemsare shown i n f i g u r e 3. The stabilizercontrolsystem i s somewhat more complex. There is a springtoprovideforcesinsteady maneuversand t h e v a r i a t i o n o f t h i s f o r c e w i t h s t i c k p o s i t i o n i s shown i n figure 4. It shouldbenoted that thespringforce varies l i n e a r l y w i t h s t i c k d e f l e c t i o n and that there i s an i n i t i a l p r e l o a d i n t h e feel springofabout 1 pound. The springpreloadforce combined w i t h s t i c k f r i c t i o n f o r c e s results i n a breakout force of about 3 t o 5 pounds.
111.11.1111111111.1 111 111 I I I 1 I I 111111 I I I I I I 1111111111111 11111 I I I II I I II 11111 I I I I 111 1 1 1 II I I 111 I II II I I 111 I I I I I- Bobweights are used t o provide additional forces when t h e a i r p l a n e is i n a c c e l e r a t e d f l i g h t . There are two bobweights,one located at t h e s t i c k andone at t h e tail, which a r e s e n s i t i v e t o b o t h normal acceleration and pitchingacceleration.Insteadyturnsor maneuverswhere t h e normal acceleration is f a i r l y s t e a d y at valuesabout l g , t h e . f o r c e s from t h e two bobweightsopposeeach other and provide a force at t h e s t i c k o f 2.6 pounds per g. In the transient portion of a maneuverwhere pitching acceleration occurs, the forces from t h e bobweights combine t o produce a force at t h e s t i c k p r o p o r t i o n a l t o p i t c h i n g a c c e l e r a t i o n o f 9.3 lb/radians/sec2.Additionalforces are providedduringstickmotion by two dampers,one located at t h e s t i c k andone at t h e tail. These dampers provide a force of 3.4 pounds per inch per second of stick deflection. A relief valveineachofthe dampers is set so that a force of 30 pounds is t h e maximum force that canbeproduced by t h e com- bined dampers. I n a d d i t i o n t o t h e f o r c e c h a r a c t e r i s t i c s p r o v i d e d i n t h e a i l e r o n and s t a b i l i z e r c o n t r o l systems,there is a nonlinear linkage in these control systems which results in a low gearing between surface and s t i c k d e f l e c t i o n n e a r n e u t r a l and increasing gearings as t h e s t i c k is deflected away from neutral.Inthelongitudinalcontrolsystemthe f e e l s p r i n g is i n t h e rear portion of the fuselage just ahead of t h e A s a r e s u l t , some s t i c k d e f l e c t i o n i s r e q u i r e d t o nonlinearlinkage.
take up the backlash before the stabilizer moves and t h i s t e n d s t o a c c e n - tuatethenonlinearity. These r e l a t i o n s h i p sa r ei l l u s t r a t e di n figure 4.
The t r i m systems i n t h e a i r p l a n e a r e u n i q u e i n that a l l three of the cockpit controls have t h e same n e u t r a l p o s i t i o n r e g a r d l e s s of t h e t r i m . The trim actuators are controlsurfacepositionrequiredfor extendablelinksinthecontrolsystems.For example, i f t h e p i l o t is holding a s t i c k d e f l e c t i o n , and thereby a c e r t a i n s t i c k f o r c e , i n o r d e r t o maintain a given airspeed and he wishes t o t r i m t h e system t o z e r o force, he must move t h e c o n t r o l s t i c k backtowardneutral as he t r i m s t h e s t a b i l i z e r t o t h e p o s i t i o n n e c e s s a r y t o h o l d t h e d e s i r e d trim speed.
The same c o n d i t i o ne x i s t si n a l l threecontrols. The trim systems a r e electronically controlled systemswhichoperatethroughtheautomatic controlamplifiers.Potentiometerslocatedonthestickgrip and on t h e l e f t console are used t o i n t r o d u c e s i g n a l s t o t h e t r i m system. The output of the trim actuators are proportional to the given potentiometer knob position. The longitudinalcontrolsystemhasan emergency trim systemwhich when o p e r a t e d c a l l s f o r t h e m a x i m u m trim actuator rate whilethe emergency switch is engaged.Thistype of t r i m system i s commonly c a l l e d a "beep" type of t r i m system.
Automatic s t a b i l i z a t i o n of t h e a i r p l a n e is providedaboutthe yaw and roll axes in both the landing and cleanconditions. The yaw damper is controlled by two independent lateral accelerometers located near thecenter of gravity. Two signals, one from eachaccelerometer,each ofwhich supplies one-half the required magnitude are transmitted through two altitudegainchangerstotheamplifiers. The a l t i t u d e gain changers increase the damper gain with increasing altitude. Signals from the two amplifier channels are fed to dual electrohydraulic actuators and result in the required surface displacement through the combined stroke of both ends of the dual actuator. A n aileron-rudder interconnect circuit is combined with the yaw damper system to prov%de rudder def tion ina roll maneuver as a function of aileron position. The rudder is used to counteract the favorable yaw produced by the ailerons. The favorable yaw decreases with increasing angle of attack; therefore, the rudder-aileron interconnect signal is passed through a stabilizer- position gain changer to attenuate the signal as the stabilizer is m in the trailing-edge-up direction. The aileron-rudder interconnect does not function in the landing condition.
The roll damping system receives its signals from two rate -os, one used for the clean condition and one for the landing condition. In the clean condition the gain between roll rate and aileron position i constant at 0.14O of total aileron per degree per second rate of roll.
In the landing condition the initial gain is 1.4O of total aileron per degree per second rate of roll. A gain changer in the landing condition reduces the gain from 100 percent to 40 percent in the first 2 inches (l/3 of full travel) of lateral stick displacement and from 40 percent to 0 percent as the stick displacement is increased from 2 to 6 inches .
(full travel) For these tests the center of gravity of the airplane was locate at 0.263; at a take-off gross weight of 26,077 pounds with the gear down. Retraction of the landing gear moves the center of gravity for- ward O.OO3c‘.
1 N ” E N T A T I O N Standard NACA photographically recording instruments, synchronized with a timer, were used in the test airplane. An NACA designed airspeed head located on a boom at the nose of the airplane was used to measure total and static pressures. Also, the head contained flow-direction vanes for measuring angle of attack and sideslip angle. The following quantities were measured and recorded: Stabilizer position Aileron position Rudder position Stick position Rudder pedal position Stick force Rudder pedal force Angle of attack S i d e s l i p angle Airspeed Altitude Three componentsof acceleration Rolling velocity and acceleration Pitching velocity and acceleration Yawing v e l o c i t y and acceleration Wing p o s i t i o n Wing s t r u t f o r c e N o c a l i b r a t i o n of t h e boom and airspeed head as i n s t a l l e d i n t h i s airplane was made. The airplanemanufacturer, however, hascalibrated a nose boom i n s t a l l a t i o n similar t o t h i s i n s t a l l a t i o n and t h i s c a l i b r a - was used t o c o r r e c t t h e measured airspeed. A p l o t of t h e c a l i b r a - t i o n t i o n is shown i n f i g u r e 5 . Also, f i g u r e 5 presents a comparison of t h i s calibration with a calibration obtained from the data presented in reference 17. Inaddition, a pointobtained from theairspeed-altimeter a t the time of thestatic-pressure jump i s presented. The recorder was consideredzero after e r r o r i n s t a t i c p r e s s u r e and t o t a l p r e s s u r e t h e jump occurred.This one datum pointappearstoagree well withthe data obtained from reference 17. The two c a l i b r a t i o n s are i n good agreementthroughoutthe Mach number range. It should be notedthat t h e c a l i b r a t i o n is p l o t t e d as a function of indicated Mach number and t h a t a d i s c o n t i n u i t y e x i s t s i n t h e c a l i b r a t i o n c u r v e s at thetime of t h e shockpassageoverthenose boom s t a t i c o r i f i c e s . The c a l i b r a t i o n i s actuallynonexistent from M = 0.96 t o 1.02.
A camera was installed in the cockpit to photograph a t a r g e t air- planethroughthewindshieldduringtrackingtests. It was notpracti- c a l t o photograph through the pilot's gunsight but the camera was bore- sighted so that tracking errors could bedetermined.
The manufacturer'svaluesofthe moments o fi n e r t i a Ix, Iy,.and .
Iz were used i nc a l c u l a t i n gc e r t a i ns t a b i l i t yd e r i v a t i v e s . These moments-of-inertiavalues were corrected for changes i n w e i g h t , d u e t o f u e l consumption.
TESTS, RESULTS, AND DISCUSSION Longitudinal Stability and Control S t a b i l i t y and c o n t r o l c h a r a c t e r i s t i c s i n s t e a d y f l i g h t . - F l i g h t tests were made t o measure the static stability throughout the speed range in the clean condition a t both 35,000 f e e t andabout 20,000 feet.
These tests wereperformed by trimming the airplane at some highsubsonic speed, and then by decreasing the speed and accelerating from some mod- erate subsonicspeed t o aboutthe m a x i m u m level-flightspeed. The speed changes were accomplishedbyvaryingtheenginethrottle. It shouldbe n noted, however, that t h e changes i n t r i m with power s e t t i n g are small and would notbeexpected t o have a s i g n i f i c a n t e f f e c t on t h e s t a b i l i z e r v a r i a t i o n s w i t h Mach number. The p i l o t a t t e m p t e d t o m a i n t a i n f l i g h t at l g throughout the tests and only those data were used except for some few c a s e s i n which t h e d a t a were c o r r e c t e d t o l g f l i g h t . These tests alsoprovided a measure of the transonic trim change. The d a t a f o r t h e two t e s t a l t i t u d e s are presented in figure 6. The d a t a show p o s i t i v e s t a b i l i t y f o r a l l Mach numbers except in the transonic speed range.
I n s t a b i l i t y is indicated from a Mach number of 0.92 t o 1.03. The s t i c k forces associated w i % h the transonic trim changes are small, on t h e order of 2 t o 3 pounds,and are considered desirably small by t h e p i l o t s .
The abrupt change i n s l o p e of the curve of stick force plotted against Mach number a t a Mach number of 0.8 i s a r e s u l t of t h e f l a t s p o t i n t h e s t i c k - t o - s t a b i l i z e r r e l a t i o n s h i p t o g e t h e r w i t h t h e s p r i n g p r e l o a d a n d s t i c k f r i c t i o n . T h i s results i n t h e f o r c e of 2$ t o 3 pounds on e i t h e r side of t r i m shown i n f i g u r e 6 ( b ) .
The same type of test was performed in the landing condition by gradually decreasing the airspeed from 180 knots to about 125 knots.
These data arepresentedinfigure 7. A s t a b l e v a r i a t i o n of horizontal t a i l position with speed is indicated although there is a slight tendency towarddecreased s t a b i l i t y a t t h e lower airspeeds. It might be noted that in the landing condition the airplane begins to undergo l i g h t b u f f e t at about an indicated airspeed of 155 knotswhich is considerablyabove t h e s t a l l i n g s p e e d of theairplane. The p i l o t s o b j e c t e d t o t h i s h i g h buffeting speed in the landing approach and f e l t that buffet could not beused a s a stall warning in this configuration.
The p i l o t s made several observations regarding the landing charac- t e r i s t i c s of theairplane. It shouldbenoted, however, that no experi- ence has beenobtainedduringcarrierlandings. The p i l o t s normally landed the airplane out of trim to avoid using the portion of t h e s t i c k - t o - s t a b i l i z e r g e a r i n g where thegearing i s low. The p i l o t s f e e l t h a t thecontinuousneed t o retrim the airplane both longitudinally and l a t - e r a l l y when the airspeed is reducedfrom 180 t o i20 knots is undesirable.
Also, t h e p i l o t s n o t e d t h a t t h e a i r p l a n e i s d i f f i c u l t t o h a n d l e d u r i n g take-offs o r landings in moderate cross windsof 10 t o 15 knotsbecause ofexcessiveheeling andweathercocking. I n t h i s p a r t i c u l a r a i r p l a n e , during the landing approach the roll stabilization system i s frequently turned off as a r e s u l t of t h e roll mon’itoring c i r c u i t when l a r g e a i l e r o n deflections are used.This is undesirableespeciallyduringanapproach in turbulent air becausethe roll stabilization system i s the system which i s most e f f e c t i v e i n damping theairplanemotions. (It was later found t h a t a malfunction of one of the gyros used for the landing con- d i t i o n was thesourceofthetrouble.) Also, t h e p i l o t s n o t e d that t h e r e s t r i c t i o n t o 220 knots airspeed with the wingup demands very careful attention during an afterburner take-off to insure that t h e wing is I lowered and locked before the airspeed is exceeded.
C h a r a c t e r i s t i c s i n a c c e l e r a t e d f l i g h t . - The maneuver characteris- t i c s of t h e a i r p l a n e weremeasuredbyperforming windup t u r n s at various a l t i t u d e s and f o r a rangeof Mach numbers. I n a l l ofthe tests at supersonicspeedsthecruisedroop was up, but at subsonicspeeds tests were made w i t h t h e c r u i s e droopboth up and down. I n most cases,the acceleration w a s i n c r e a s e d i n t h e windup t u r n s u n t i l moderate b u f f e t w a s encountered. Some t e s t s were a l s o made t o determinethecharacter- i s t i c s i n r a p i d p u l l - u p s and t u r n e n t r i e s .
The s t i c k f o r c e , s t i c k p o s i t i o n , and s t a b i l i z e r p o s i t i o n as a func- t i o n ofnormal a c c e l e r a t i o n i n windup turns at a l t i t u d e s of about 30,000 and 35,000 f e e t f o r two calibrated Mach numbers a r e p r e s e n t e d i n fi,o;ure 8. The data of f i g u r e 8 aretypicalofthedataobtainedduring t h e f l i g h t program. The v a r i a t i o n of stabilizeranglewithaccelera- t i o n i s s t a b l e and l i n e a r i n a l l cases. The stick-force and s t i c k - position curves reflect the nonlinearity of the control systemand t h e e f f e c t offuselagebending. The breakoutforcerequiredto ovelrcome t h e s t i c k f r i c t i o n and spring preload together with the forces resulting from the very low gearing between s t i c k and s t a b i l i z e r n e a r n e u t r a l requires a s t i c k f o r c e of about 3 t o 5 pounds t o move t h e s t a b i l i z e r .
These f o r c e s c a u s e t h e i n i t i a l f o r c e p e r g for values ofnormal accel- eration up t o about 2g t o exceed the limits s p e c i f i e d i n t h e r e q u i r e - ments of reference 11. The forceper g forvalues of g i n excessof 2g are well within the required limits. The data withcruise droop up indicatethe same trends as those for the cruise-droop-downcase, and t h e s t a b i l i z e r a n g l e p e r g i s slightly less for the cruise-droop-up condition. Windup turns performed at a n a l t i t u d e of 20,000 f e e t w i t h thecruisedroop down exhibit the same c h a r a c t e r i s t i c s as thoseobtained a t 35,500 f e e t . The stabilizerangleperg, however, i s decreased because of the increase in dynamic pressure. The l o w e s t a l t i t u d e f o r which t e s t data are presented was 14,400 f e e t at a Mach number of 0.9.
These data are presented in figure 9 and show that t h e s t a b i l i z e r v a r i a -
t i o n w i t h normal acceleration i s approximately linear up t o t h e h i g h e s t -
value of g reached. The stabilizerangleperg, however, i s decreased r e l a t i v e t o t h e o t h e r Mach numbers and a l t i t u d e s . The curvesof s t i c k i f any, force and s t i c kp o s i t i o na r e of s p e c i a li n t e r e s t . Very l i t t l e , s t i c k motion i s r e q u i r e d t o move t h e s t a b i l i z e r at thehighervalues of normal acceleration but the forces required are almost linear and r e f l e c t t h e f o r c e r e s u l t i n g from thenormal-accelerationbobweights.
w a s overly sensitive at t h i s Mach num- The p i l o t s f e l t t h a t t h e a i r p l a n e ber and a l t i t u d e , b u t f o r slow steady maneuvers t h i s c h a r a c t e r i s t i c was nottooobjectionable.Ingeneral,thepilots wereof theopinion that the longitudinal control i s too insensitive near t r i m f o r a l l regions
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of f l i g h t w i t h t h e e x c e p t i o n ofindicatedairspeeds i n excess of 500 knots.Thiscausesthesystem t o be p a r t i c u l a r l y annoying w h i l e tracking or during the beginning of the landing flare from a t r h n e d condition. The nonlinearvariation of s t i c k - t o - s t a b i l i z e rr e l a t i o n is responsible for this deficiency and it is f e l t that a more l i n e a r con- t r o l system, e s p e c i a l l y f o r moderate control displacements, wouldbe an -improvement.
A summary p l o t of t h e s t a b i l i z e r angle per g i n a c c e l e r a t e d maneu- vers is shown i n figure 10. The data of figures 8 and 9 togetherwith a l l of t h e measured d a t a i n a c c e l e r a t e d maneuvers are p r e s e n t e d i n t h i s figure. A t an a l t i t u d e of35,000 feet thestabilizerangleper g decreases somewhat abruptly from about 3.5 t o about 2.8 in the range of Mc from 0.92 t o 0.97 and thenincreasesrapidly as supersonic speedsareattained,reaching a m a x h u m of about 5.3 a t Mc = 1.1.
Above t h i s Mach number and up t o about M, = 1.45 t h es t a b i l i z e ra n g l e per g decreases until a valueaboutthe same o r s l i g h t l y l e s s t h a n t h a t forthesubsonicconditionexists.Puttingthecruisedroop up at sub- sonicspeeds at 35,000 feetcauses a slight decrease i n s t a b i l i z e r a n g l e required. A t the lower a l t i t u d e of 20,000 f e e t ,t h es t a b i l i z e ra n g l e per g decreasedwithincreasing Mach number fromabout2.5 a t Mc = 0.68 t o 1.65 a t Mc = 0.865. The minimum value of 1.5 was obtained a t Mc = 0.9 a t an a l t i t u d e of approximately 14,000 f e e t .
A number of f l i g h t t e s t s were made of rapid pull-ups and t u r n entries to obtain pilot opinions of t h e 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 airplane undertheseconditions.Typicaltimehistories of pull-up maneuvers arepresentedinfigure 11. Thesemaneuverswereof particu- lar i n t e r e s t because of theiongitudinalfeelsystem. The p i l o t s f e l t that t h e f o r c e c h a r a c t e r i s t i c s i n r a p i d maneuverswere very good. The f o r c e d u r i n g t h e i n i t i a l p a r t of the maneuvers was somewhat higherthan in steady turns. There was a tendency f o r t h e p i l o t s t o overshootthe desired acceleration level when rapid turns to large accelerations were made. However, t h i s tendency was believedto be due t o t h e n o n l i n e a r gearing and the decrease in apparent stick-fixed stability at higher g l e v e l sr a t h e rt h a nt ot h ef o r c ec h a r a c t e r i s t i c s . A l s o , t h e p i l o t s f e l t that there w a s l i t t l e tendencytoward p i l o t induced oscillations and that t h e f e e l system d i d n o t r e s t r i c t t h e maneuvering c a p a b i l i t i e s of the airplane.
Some tests were made t o measure t h e maneuver c h a r a c t e r i s t i c s i n thelanding-approachconfiguration f o r a rangeofairspeeds from 200 knots down t o 140 knots. These data are presentedinfigure 12.
About t h e same trends of s t a b i l i z e r p o s i t i o n and f o r c e c h a r a c t e r i s t i c s areexhibitedinthelandingcondition as inthecleancondition. In the landing condition the airplane begins to buffet at very small incre- ments of gabove l g and it w a s d i f f i c u l t t o m a i n t a i n t h e t u r n at any
-
given g l e v e l . T h i s c o n d i t i o n r e s u l t e d i n t h e amount of scatterobtained i n t h e d a t a . The f o r c e p e r g inthelandingcondition is somewhat large,
on the order of 15 pounds per g. The s t a b i l i z e r a n g l e p e r g increases -
fromabout 4 . 5 O per g a t 197 knots t o about loo t o 1 2 ' per g at 140 knots.
E f f e c t s of fuselagebending.- As has been noted, the data of fig- ure 9 i n d i c a t e t h a t t h e r e l a t i o n between t h e s t i c k and s t a b i l i z e r motion i s adversely affected by normal acceleration to such an extent that t h e r e is a l a r g e d e c r e a s e i n a p p a r e n t s t i c k - f i x e d s t a b i l i t y ; t h a t is, at Mach numbers near 0.9, t h e v a r i a t i o n o f s t i c k p o s i t i o n w i t h normal a c c e l e r a t i o n i n d i c a t e s t h a t t h e a i r p l a n e is n e u t r a l l y s t a b l e , whereas t h e v a r i a t i o n of s t a b i l i z e r a n g l e w i t h a c c e l e r a t i o n shows t h a t t h e air- planehas a sizeable margin of s t a b i l i t y . These data i n d i c a t e that t h e longitudinal control system i s affected by loading on some portions of theairplanestructureorcontrolsystem.Inanefforttoisolatethe p a r t s of thecontrolsystem which are affected, instruments were i n s t a l l e d t o measure t h e motion of v a r i o u s p a r t s of the longitudinal are shown schematically i n controlsystem. The locations of t h e s ep a r t s f i g u r e 13. I no r d e rt o measure t h ee f f e c t so f normal acceleration on the longitudinal control system, windup t u r n s i d e n t i c a l t o t h o s e d e s c r i b e d in the section entitled "Charactersitics in Accelerated Flight" were made a t d i f f e r e n t Mach numbers and a l t i t u d e s , and the data obtained from these t e s t sa r ep r e s e n t e di nf i g u r e 14. The p o s i t i o n of thestick,walking beam, structuralfeedbacklinkage, and r i g h t s t a b i l i z e r i n t e r m s of an equivalent stick position are shown as a function of normal acceleration; that is, thevariouslinkages were calibrated in terms of s t i c k a n g l e so t h a t on t h e groundunder no load a l l of the curves would coincide.
The difference between the curves in flight indicates the deformation occurring a t various points in the control system in terms of t h e s t i c k angle whichwould be r e q u i r e d t o produce t h i s motionunder a no-load condition. The r e s u l t s of t h e s et e s t si n d i c a t et h a ta l m o s t a l l of the loading effects due to acceleration occur between t h e s t i c k and t h e walking beam. There a r es l i g h td i f f e r e n c e s between t h ep o s i t i o n of t h e walking beam and the structural feedback linkage but these effects are small compared with the differences between t h a t of t h e s t i c k andwalking beam. A comparisonbetween thewalking beam and t h e motion of t h e sta- b i l i z e r a l s o i n d i c a t e s o n l y s l i g h t d i f f e r e n c e s whichcanprobablybe accountedforintheaccuracyoftheinstrumentation. It mightbenoted that only the output of the structural feedback linkage was measuredand t h a t some compensation f o r s t r u c t u r a l motioncouldbeoccurringwhich would notbemeasured by t h e i n s t r u m e n t a t i o n i n s t a l l e d f o r t h e s e t e s t s .
A p l o t of t h e d i f f e r e n c e between t h e s t i c k motionused t o o b t a i n a given g and t h e s t i c k motionwhich would have been required t o produce t h e same amount of s t a b i l i z e r d e f l e c t i o n on t h e ground i s shown i n f i g u r e 15.
Data are presented for a l l t h e t e s t c o n d i t i o n s of Mach number and dynamic pressure. The r e s u l t si n d i c a t et h a ta c c e l e r a t i o nl o a d s on theairplane cause the longitudinal control system to deflect the stabilizer an amount equivalenttoabout 0.90 of s t i c k motionperg.Thisplotalsoindicates .
that the amount of deflection is almost independent of dynamic pressure or Mach number, at least for the range of the test conditions. A s shown by the previous data on windup turns, the most serious effects of the undesirable control-system motion occur at Mach numbers around 0.9 where the stabilizer angle per g is the smallest. The data obtained from these tests indicate that the control-system movement due to def mation results from bending of the forward portion of the fuselage brought about by inertia loading during accelerated maneuvers.
It might be noted that the airplane manufacturer has redesigned longitudinal control system to account for the effects of fuselage be The change to the control system has not as yet been tested by the NASA, but flight tests by the manufacturer indicate that the linkage change alleviated the problem.
In the performance of maneuvers to high acceleration some marked changes in the aerodynamic stability characteristics were found to exist at the higher values of acceleration at Mach numbers of about 1.0, 1.1, and 1.2. These decreases in stick-fixed stability can be 14 and tend to aggravate the structural seen in the data in figure deformation effects at the higher values of acceleration.
Stability derivatives as determined from _ _ _ dynamic stability tests.- - ." ~~ -~ " . . - - - - - - ~ ~ - ~ - " _ _ ~ -~ In order to measure the dynamic stability characteristics of the airpla pulse stabilizer inputs were imposed on the airplane for the Mach numb range of the airplane at an altitude of approximately 35,000 feet. The resulting period, time to damp to one-half amplitude, and damping ratio were obtained from the short-period oscillation. These data are pre- sented in figure 16.
The period changed from about 2.3 seconds at M = 0.8 to about 1 . 5 seconds at M = 0.92 and then changes slowly to about 1.0 second at M = 1.4. The time to damp to one-half amplitude varies from about 1.25 seconds at low Mach numbers to about 0.8 second at M = 1.44. The sharply from about 0.20 at M = 0.8 resulting damping ratio decreases to about 0.17 at M = 0.92, reflecting the large change in stability 0.9. The damping ratio is about constant at a at Mach numbers around value of about 0.14 from about M = 1.0 to 1.4.
The pilots considered the damping of the short-period longitudinal oscillation to be low and less than desired. The poor damping did not materially affect the performance of the airplane during general flying which involved only gradual maneuvers. However, the lack of good damping does result in more work during such tasks as tracking and i particularly bothersome while tracking a maneuvering target. Some brief tests regarding the tracking capabilities of the airplane are discusse subsequently.
I.
.
The v a r i a t i o n o f . s t a b i l i t y w i t h Mach number as shown by the param- e t e r CnLL is presentedin figure 17. Thisparameter was obtained from 1) the expression the period and damping data by using
cma =
Also, shown i n figure 17 i s t h e summation of the rotary derivatives
+ %. These data were obtained from the formula
cmq
cmq + C% = si-.(-) + CNa g j
The lift-curve slope of t h e a i r p l a n e w a s a l s o measuredfrom t h e a i r p l a n e short-period oscillation by measuring t h e normal acceleration and angle of a t t a c k d u r i n g a n o s c i l l a t i o n i n p i t c h . The followingequation was used to obtain the values shown i n figure 18: The lift-curve slope appears to reach a maximum of 4.5 per radian at M = 0.92. The slopedecreasesgradually above M = 0.92 t o about 3 .O perradian at M = 1.4. The v a r i a t i o n of static margin dCm/dCn with Mach number as obtained from the measured values of C and C
ma Na
i s p r e s e n t e d i n f i g u r e 19. The airplane has a s t a t i c marginofabout 17.5 percent c' a t Mach numbers from 0.76 t o 0.85 and then changes rapidlytoabout 30 percent at M = 0.96. A s t h e Mach number increases, t h e s t a t i c margingraduallyincreases t o about 33 percent E at M = 1.44.
D i r e c t i o n a l S t a b i l i t y and Control S t a b i l i t y and c o n t r o l c h a r a c t e r i s t i c s i n s i d e s l i p . - S i d e s l i p d a t a were obtained in the clean condition at a l t i t u d e s of approximately 35,000 f e e t and20,000 feet. Also, s i d e s l i p d a t a were obtainedforthe landing configuration at airspeeds of200and150knots at 8,500 feet.
The maneuverswere made a t nearlyconstantvelocity. The rudder was used t o i n c r e a s e s i d e s l i p i n one d i r e c t i o n u n t i l a maximum deflection w a s reached; then, the controls were r e t u r n e d t o n e u t r a l and t h e same procedure was usedintheotherdirection.Sideslip data at several t e s t a l t i t u d e s and Mach numbers are presented in figure 20. The data consist of plots of control-surface positions for the aileron, rudder, .
and horizontal t a i l and t h e p i l o t s ' c o n t r o l f o r c e s n e c e s s a r y t o hold these positions as a function of sideslipangle. The a i l e r o n and rudder control-surface positions varied linearly with sideslip and were i n t h e stable direction throughout the Mach number range of the tests. The stabilizer position did not vary with sideslip. The rudder pedal force was l i n e a r w i t h s i d e s l i p a n g l e and t h e a i l e r o n f o r c e r e f l e c t e d t h e non- l i n e a r r e l a t i o n s h i p between s t i c k and aileron deflection. The maximum a i l e r o n f o r c e w a s generally less than 10 pounds. The maximum peaal force w a s between 150 and 200 pounds f o r maximum rudder deflection.
P l o t s of - d6r and - d6a forvarious Mach numbers at a l t i t u d e s of
dP dP approximately 35,000 f e e t are presented i n f i g u r e 21. The i n c r e a s e i n dEa - above Mc = 1.0 indicates a decreaseinaileroneffectiveness, dP as shown i n a subsequentsection, and a possibleincrease i n t h e r o l l i n g dfjr
moment due t os i d e s l i p . The parameter - alsoincreases a t Mach
dP numbers above Mc = 1.0. The increaseinthisparameter i s due mainly t o a large reduction in rudder effectiveness at supersonicspeeds.
Sideslip data for the landing condition are presented in figure 22.
Data are presented for three different airspeeds which represent a spreadinnormal-forcecoefficient from 0.44 t o 0.99. Inthelanding configurationtheavailableruddertravel i s i n c r e a s e dt o +lTO. The control-surface positions show a l i n e a r v a r i a t i o n w i t h s i d e s l i p f o r moderateangles of s i d e s l i p . The rudderforcehas a l i n e a rv a r i a t i o n w i t h s i d e s l i p and t h e a i l e r o n f o r c e r e f l e c t s t h e n o n l i n e a r v a r i a t i o n of ailerondeflectionwithstickdisplacement. The amount of a i l e r o n and rudder deflection per degree of sideslip in the landing configuration i s larger than t h a t shown f o r t h e c l e a n c o n d i t i o n ( f i g . 20) a t the lowest Mach numbers. A t thehighestnormal-forcecoefficient ( f i g . 2 2 ( c ) ) t h e r e a p p e a r s t o b e some decrease in the directional sta- b i l i t y and t h e r o l l i n g moment due t o s i d e s l i p is somewhat greater as evidenced by t h e v a r i a t i o n of rudder angle and a i l e r o n angle with side- s l i p . Also, t h e r e i s anincreaseinpitching moment due t o s i d e s l i p as shown by t h e v a r i a t i o n of stabilizerangle.Thisconditiondidnot e x i s t a t thelowernormal-forcecoefficients.Although figure 22 does not show t h a t t h e maximum aileron deflection is reached, the pilots noted that maximum aileron deflection was reached before maximum rudder deflection.
Rollperformance.-Although most of theregimes of f l i g h t of t h e airplane have beencovered, no d e t a i l e d f l i g h t s t u d y has yet been made of t h e r o l l i n g performanceoftheairplane.Results from a preliminary study of r o l l performance based on data obtained from Chance Vought Aircraft, Inc. and some f l i g h t d a t a f r o m t h e Langley Flight Research Divisionarepresented. It shouldbenoted that thedatapresented P 1-7 W A data show s l i g h t l y h i g h e r r o l l i n g v e l o c i t i e s . The d a t a of f i g u r e 24 show t h a t below M = 1.0 over most of theusablerangeofaltitudethe test airplane can meet the proposed r o l l s p e c i f i c a t i o n o f goo i n 1 second.
S t a b i l i t y d e r i v a t i v e s and other measurements determinedfrom dynamic s t a b i l i t y tests.- The dynamic lateral d i r e c t i o n a l s t a b i l i t y c h a s a c t e r i s - t i c s were obtained by making pulse-type inputs with the rudder and then measuring theensuingoscillations. These t e s t s were performed at an a l t i t u d e of about 35,000 feet at various Mach numbers w i t h t h e s t a b i l i - zationsystems on and o f f . The period, time t o damp t o one-halfampli- tude, and t h e damping r a t i o as a function of Mach number obtained from t h e s e t e s t s a r e p r e s e n t e d i n f i g u r e 25. The periodforthecase of s t a b i l i z a t i o n system on varies from 2.6 seconds at M = 0.76 t o about 1.8 seconds at M = 0.95. From M = 0.95 t o M = 1.3 theperiod is almostconstant at about 1.75 seconds.Thereappears t o be a tendency f o r t h e p e r i o d t o i n c r e a s e s l i g h t l y a s t h e Mach number is increased beyond 1.3 b u t t h e r e a r e i n s u f f i c i e n t data t o e s t a b l i s h t h i s t r e n d .
The pilot opinion of t h e dampingof t h e l a t e r a l d i r e c t i o n a l o s c i l l a t i o n indicated that the damping was adequate for large amplitude disturbances but the damping was consideredpoor when small disturbancesorchanges i n trim occur. The change i n damping withamplitude may result from backlash in the yaw damping systemwhich has been improved i n l a t e r versions of t h i s a i r p l a n e . The periodforthecase of t h e s t a b i l i z a - t i o n systemoffexhibitsthe same trends as that f o r t h e s t a b i l i z a t i o n system on, the period being about 0.1 t o 0.2 secondlonger i n most cases. The time t o damp t o one-halfamplitude and the damping r a t i o show the marked e f f e c t s of thestabilizationsystems. With t h e stabi- l i z a t i o n system on the time t o damp t o one-halfamplitude 16 f a i r l y constantatabout 1 second up t o M = 1.3 and the damping r a t i o v a r i e s from about 0.27 at M = 0.76 t o about 0.18 at M = 1 . 3 . Here again there appears to be a trend toward increasedtime t o damg t o one-half amplitude a t Mach numbers above 1 . 3 . The stabilization-system-offcase shows the time to damp t o one-halfamplitudevaries fromabout 2.1 sec- onds a t M = 0.82 t o about 1.5 seconds a t M = 1.37 and t h e damping r a t i o is aboutconstant a t 0.13 t o 0.15. The pilotsconsideredthe damping of t h e l a t e r a l d i r e c t i o n a l o s c i l l a t i o n t o b e poor with the s t a b i l i z a t i o n system o f f .
The r o l l - t o - s i d e s l i p r a t i o s measured during the lateral directional o s c i l l a t i o n sa r ep r e s e n t e di n figure 26. The s t a b i l i z a t i o n system decreases the roll-to-sideslip ratio at a l l Mach numbers throughout the speedrange. The percentdecrease is greatest at Mach numbers from about 0.73 t o 1.13. Above M = 1.13 t h es t a b i l i z a t i o n system has less e f f e c t on t h e r o l l - t o - s i d e s l i p r a t i o b u t t h e r a t i o is still less than with the stabilization system o f f . The p l o t of the reciprocal of the
cycles t o damp t o one-halfamplitude as a flmction of theparameter A
ve is shown i n figure 27. The requirements as s e t f o r t h ' i n r e f e r e n c e 1 1 . .
- ..
f o r t h e s t a b i l i z a t i o n system both on and o f f a r e a l s o shown i n t h e s e p l o t s . The airplanemeetstherequirementsin a l l casesboth w i t h the s t a b i l i z a t i o n system on and o f f . The p i l o t s f e l t that the r o l l - t o - ..
s i d e s l i p r a t i o was highalthoughnottooobjectionable. They f e l t that these high ratios would notbe a serious factor during most f l i g h t con- d i t i o n s andwouldonlybenoticed i n maneuvers made s p e c i f i c a l l y t o measure t h i s c h a r a c t e r i s t i c .
The static directional-stability parameter was determined
CnB
from the period and damping data by the following expression: The values of Cn obtainedin this manner f o r the case of s t a b i l i z a - B t i o n systemoffarepresentedinfigure 28. The data indicate that varies fromabout 0.14 per radian at M = 0.83 t o 0.185 per radian CnP a t M = 0.93 and thendecreasesgradually t o about 0.10 perradian at M = 1.3. The directional-stabilityparameterdecreasesto a low value of about 0.08 per radian a t Mach numbers around 1.4.
The v a r i a t i o n of side-force coefficient with sideslip angle cyB was determinedfrom the followingexpression: The r e s u l t so b t a i n e da r es h o wi nf i g u r e 29. The side-forcecoefficient C has very l i t t l e v a r i a t i o n w i t h Mach number, remaining at a value of YP about -0.8 throughout the Mach number range. Only thosedataforthe s t a b i l i z a t i o n systemoffarepresented.
Trim Systems The t e s t a i r p l a n e u t i l i z e s a positional type of servocontrol in the longitudinal t r i m system; that is, t h e p i l o t p o s i t i o n s awheel on t h e s t i c k which c a l l s f o r a givenstabilizerdisplacement. This is i n c o n t r a s t t o a conventional "beep"system i n which the t r i m actuator moves a t a constant rate and stops movingwhen t h e p i l o t r e l e a s e s t h e trim control. Also, sincethestick has t h e same n e u t r a l p o s i t i o n f o r a l l c o n d i t i o n s of f l i g h t , t h e p i l o t i s r e q u i r e d t o move t h e s t i c k back toward n e u t r a l as theairplane i s trimmed. The main pilotobjections t o t h e trim system would seem t o result from t h e . f a c t that t h e f i n a l trim position of t h e s t a b i l i z e r is notreached when t h e p i l o t s t o p s the motionof the trim wheel.Becauseof this timedelay i n stabilizer motionand t h e i n a b i l i t y t o a n t i c i p a t e t h e final r e s u l t of the trim c o r r e c t i o n , t h e p i l o t r e s o r t s t o making minuteadjustments of t h e t r i m wheel. As a result, t h e p i l o t is required to use a great deal ofcon- centration not normally associated with a conventional trim system.
The trim procedure is f u r t h e r complicated by the nonlinear stick-to- s t a b i l i z e r r e l a t i o n s h i p .
Time h i s t o r i e s which i l l u s t r a t e t h e p i l o t ' s trim procedure are shown i nf i g u r e 3 0 . The first case(fig. 3O(a)) is one i n which the pilot attempted to trim t h e a i r p l a n e r a p i d l y i n a f l i g h t regime where theairplane is s e n s i t i v e t o small controlmotions. The figure shows thelargeoscillations that r e s u l t .I nt h e secondcase (fig.3O(b)) t h e p i l o t used a t r i m procedure more t y p i c a l of the normal technique used.Inthiscase, no largetriminputsareused and t h e r a t e of t r i m is minimized. However, evenunder theseconditionstheairplaneoscil- latesinpitch.Inbothcases,thetimehistory of s t i c k p o s i t i o n i n d i - c a t e s t h e p i l o t s moved t h e s t i c k i n a s e r i e s of steps.
On s e v e r a l f l i g h t s , t h e p i l o t s used the emergency t r i m system which i s a"beep" type of trim control. All t h e p i l o t s f e l t that t h i s system may bean improvement over thepresentsystem. However, since the control for the emergency system i s located on the left console and not on t h e s t i c k , it is hard t o make a comparison.
In the landing configuration the pilots found it d i f f i c u l t t o make thelarge trim changes requiredduringthelandingapproach.This com- ment is a r e s u l t of t h e l i m i t e d r a t e of t r i m actuation available in the system. The pilotsnotedthatthey had t o w a i t severalseconds before being able to determine how much trim had beenapplied and, as a result, either overshot or undershot the desired t r i m position.
The p i l o t s c o n s i d e r t h e l a t e r a l trim systempoorbecause of the d i f f i c u l t y r e q u i r e d t o trim precisely.This t r i m system is a l s o com- p l i c a t e d by thenonlineargearing between t h e s t i c k and ailerons. The l a t e r a l trim system is p a r t i c u l a r l y bothersome i n maneuvers such as tracking or when small d i r e c t i o n a l t r i m changesoccursuch as i n t h e transonic speedrange. These d i r e c t i o n a l t r i m changescausedispropor- t i o n a t e l y l a r g e l a t e r a l t r i m changes because of the large roll-to- s i d e s l i p r a t i o s and t h e e f f e c t s of thenonlineargearing. The p i l o t s considered the directional trim system s a t i s f a c t o r y and easy t o use.
The stick-force changeswhichoccur when making changes i n power, dive flap position, cruise-droop position, and wing or gear position a r e considereddesirably low. There a r e some r a t h e r l a r g e trim changes r e a c h t h e recom- when making afterburner take-offs and attempting .to mended climbspeed as r a p i d l y as possible. A l a r g e d i r e c t i o n a l t r i m
changeoccurs when changing a l t i t u d e from sea l e v e l t o about 35,000 feet. -
change is i n i t i a l l y If a r a p i d climb t o h i g h a l t i t u d e is made, t h e trim not as large. The f i n a l trim change, however, i s t h e same i f thehigh a l t i t u d e is maintainedfor any lengthof time. This trim change is common t o t h i s a i r p l a n e and is thought t o be a r e s u l t of contraction as they are exposed t o of v a r i o u s p a r t s of the rudder control system the colder a i r at high altitudes.
Some Brief Measurements of the Formation and Tracking Performance of the Test A i r p l a n e The formation flight characteristics appear good i n t h e rangeof
f l i g h t c o n d i t i o n s t e s t e d - that is, at Mach numbers about 0.9 at alti-
tudes from 10,000 t o 35,000 feet. The trackingaccuracyoftheairplane appeared t o be adversely affected by t h e poor damping of the longitudi- n a l and lateral o s c i l l a t i o n s . These opinions are based on some b r i e f tests of tracking a subsonic airplane a t an a l t i t u d e o f 35,000 f e e t w i t h t h et a r g e ta i r p l a n ef l y i n g a t M = 0.8 and t h e t e s t a i r p l a n ef l y i n g at M = 0.8 and M = 1.2. The averagestandarddeviationfortheflights made at subsonicspeeds was 3.3 mils in both azimuth and e l e v a t i o n i n a steady t a i l chase.Thesevaluesincreasedto 5.0 m i l s i n azimuthand 6.2 m i l s i n e l e v a t i o n i n t r a c k i n g a t a r g e t maneuvering a t steady g.
A t thesupersonicspeedof M = 1 . 2 i n a steady t a i l chase,thestandard deviation measured w a s 2.3 mils i n azimuthand 2.1 mils i n e l e v a t i o n . No t e s t s were made i n maneuvering f l i g h t a t supersonicspeeds. The standard deviation values for the test airplane may be compared with those for a typicalstraight-wingsubsonicairplanewhich is considered t o have good t r a c k i n gc h a r a c t e r i s t i c s . The standarddeviationvaluesforthesubsonic airplane are 1.7 m i l s i n azimuthand2.2 m i l s i n e l e v a t i o n i n a steady t a i l chaseand 3.8 m i l s i n e l e v a t i o n i n s t e a d y t u r n s . It canbeseen t h a t t h e t r a c k i n g c h a r a c t e r i s t i c s o f t h e test airplane are somewhat infe- r i o r t o t h o s e o f t h e s u b s o n i c a i r p l a n e b u t t h e t a c t i c a l u s e of t h e air- plane and the type of weapons t o beused would have t o be considered before making any definite conclusions regarding the tracking performance.
MeasurementsofLoads i n t h e Variable-Incidence-Wing S t r u t During Operationofthe Wing The t e s t a i r p l a n e has a two-position variable-incidence wing which is s e t a t - 1 ' f o r normal f l i g h t c o n d i t i o n s and r a i s e d t o 7' i n t h e landingcondition. The wing i s operatedbya.singlehydraulicstrut ( f i g . 1) which is capable of exertingabout 2,000 pounds f o r c e i n t h e .
.do&cycle.mightoperations of theairplane have indicated that the
force available at t h e s t r u t is rtkrginal in the down cycle. Yl an e f f o r t
t o e s t a b l i s h t h e l o a d s on the strut during the operation of jhe variable- incidence wing, a s t r a i n gage was i n s t a l l e d at the base of t h e s t r u t .
The forces measuredby t h i s i n s t a l l a t i o n a r e e s t i m a t e d t o b e accurate t o w i t h i n flOO pounds.
Operation of the variable-incidence w i n g i n several conditions is shown i n f i g u r e s 3 l t o 34, and operation of theleading-edge droop t o the cruise and landing positions with the wing i n t h e c l e a n p o s i t i o n is shown i n figures 35 and 36. Finally,operation of the wing with t h e drooplocked inthelandingcondition is shown i n figure 37. The r e s u l t s of t h e s e t e s t s i n d i c a t e d that it is necessary t o maintain the leading-edgedroop in the cruise position during the down cycle t o keep thestrutloadswithinthecapabilities of thehydraulicstrut. The time history of figure 34 shows that the loads in the strut reach about 2,200 pounds with the cruise droopup. The load is decreasedbyabout 300 pounds when the droop is i n t h e c r u i s e p o s i t i o n ( f i g . 35). Because of the large effect ofdroop position on t h e s t r u t l o a d s , t e s t s were made w i t h the drooplocked in the landing condition throughout the wing cycle. These t e s t s ( f i g s . 36 and 37) showed t h a tt h es t r u tl o a d sa r e decreasedbyabout1,300 pounds when the droop is d e f l e c t e d t o t h e landing position and that t h e s t r u t load during a wing-down cycle does notexceed TOO pounds.
I n o r d e r t o circumvent the problem described previously, the manu- facturer has redesigned the hydraulic actuating strut to increase the output force of t h e s t r u t i n b o t h t h e wing-up and wing-down cycles.
CONCLUDING REMARKS F l i g h t - t e s t measurementshavebeen made of t h e f l y i n g q u a l i t i e s and some of t h e s t a b i l i t y d e r i v a t i v e s of a supersonic fighter airplane.
I n addition, pilot opinion of variousaspects of the handling qualities is presented. The f l i g h t t e s t s cover a range of Mach numbejrs up t o 1.5 and an altitude range from s e a l e v e l t o 35,000 f e e t .
The damping of the short-period longitudinal odcillation is low and together with somewhat poor damBing of small amplitude l a t e r a l o s c i l - l a t i o n s resulb i n r e l a t i v e l y poor tracking performance of the airplane at subsonicspeeds. The airplaneintheopinion of t h e p i l o t s has high roll-to-yawratios; however, the airplane meets the roll-to-yaw specifi- cations of thepresentflying-qualitiesrequirements. The longitudinal f e e l system is considered good by the pilots, but some of the longitudinal- control-systemcharacteristicsresult i n poorhandlingqualities.In particular, the nonlinear relationship between t h e s t i c k and s t a b i l i z e r . . .. .
. . .. .. - results in the airplane being tod..iqsensitive for .all indicated airspeed up to about 500 knots. Even above these airspeeds, it appears as though a more linear stick-to-stabilizer gearing would be an improvement. In .
accelerated maneuvers, structural deformation of the airplane results in motion of the stabilizer without a corresponding motion of the stick.
This motion causes the apparent stick-fixed stability as measured by the stick position to become less stable and in some flight conditions, where' the stabilizer angle per g is U, the airplane stability varies from neutral to unstable. The pilots dislike the longitudinal trim system because of the difficulty experienced when w i n g to trim pre- cisely and the overshoot which occurs when making large or rapid trim corrections.
Langley Research Center, National Aeronautics and Space Adminiskration, Langley Field, Va., Ma;y 19, 1958.
, REFEREDICES 1. Anon.: Model X F 8 U - 1 Airplane,Phase I, N a v y PreliminaryEvaluation Report No. 1, SpecialReport.Project TED No. PTR AC-24101.1, Flight Test Div., U. S. Naval A i r Test Center (Patwent River, Md.), J ~ Y 6, 1955.
2. Anon.: Model X F 8 U - 1 Airplane, Naval PreliminaryEvaluation,Phase I, LetterReport No. 2. Project TED No. FTR AC-24101.1, F l i g h t Test Div . , U . S. Naval A i r Test Center (Patuxent River, Ma. ) , J u l y 14,
1.955 -
3. Anon.: Model X F 8 U - 1 Airplane,Phase 11, Navy PreliminaryEvaluation, Report No. 1, SpecialReport.Project TED No. PTR AC-24101.1, FlightTest Div., U. S. Naval A i r TestCenter(PatwentRiver, Md.), Oct. 13, 1955.
4. Anon.: Model X F 8 U - 1 Airplane,PreliminaryEvaluation,Phase 11, LetterReport No. 2.Project T E D No. PTR AC-24101.1, F l i g h t Test Div., U. S. Naval A i r T e s t Center(PatuxentRiver, Md.) , Oct . 27, 5 . Anon. : Phase IV Navy PreliminaryEvaluation of t h e Model XF8U-l/F8LJ-l Airplanes,Report No. 1. Project T E D No. FTR AC-24101.1, F l i g h t TestDiv., U. S. Naval A i r TestCenter(PatwentRiver, Md.), J u l y 6, 1956.
6. Anon. : Phase IV N a v y PreliminaryEvaluation of t h e Model XF8U-l/F8LJ-l Airplanes,Report No. 2.Project TED No. PTR AC-24101.1, F l i g h t TestDiv., U. S. Naval A i r Test Center(PatuxentRiver, Md.), Aug. 6, 1956.
7. Anon.: S t a b i l i t y and Control Trials of t h e Model F8U-1 Airplane, LetterReport No. 1, PreliminaryEvaluationPhase.Project TED No. BIS 21210, F l i g h t Test Div., U. S. Naval A i r T e s t Center (PatuxentRiver, Md.) , Oct. 1, 1956.
8. Anon.: S e r v i c eS u i t a b i l i t y Trials of t h e Model F8U-1 Airplane, Report No. 1, PreliminaryEvaluationPhase.Project TED No. BIS 2L210, ServiceTestDiv., U. S. Naval A i r TestCenter(Patwent River, Md. ) , Oct . 1, 1956.
9. Anon.: Armament Trials of Model F8U-1 Airplane(Preliminary Evaluation Phase),Report No. 1. Project TED No. BIS 21210, Armament Test Div., U. S. Naval A i r Test Center(PatuxentRiver, Md.), Nov. 1, 10. Anon.: Fleet Introduction Program of the Model F8U-1 Airplane, Report No. 1, Final Report. Project TED No. PTR AC-24101.3, Service Test Div., U. S. Naval Air Test Center (Patuxent River, Md.), Apr. 4, 1957.
11. Anon.: Flying Qualities of Piloted Airplanes. Military Specifica- tion, MIL-F-8785 (ASG) , Sept. 1, 1954.
12. Klinai., Walter J.: Spinning and Related Problems at High Angles of Attack for High-speed Airplanes. NACA RM L55L23a, 1956.
13. Klinar, Walter J.: A Study by Means of a Dynamic-Model Investiga- tion of the Use of Canard Surfaces as an Aid in Recovering From Spins and as a Means for Preventing Directional Divergence Near the Stall. NACA RM L56B23, 1956.
14. Boisseau, PeterC.: Low-Speed Roll Effectiveness of a Differentially Deflected Horizontal-Tail Surface on a 420 Swept-Wing Model. NACA RM ~ 5 6 ~ 0 3 , 1956.
15. Paulson, John W., and Boisseau, Peter C.: Low-Speed Investigation of the Effect of Small Canard Surfaces on the Directional Stability of a Sweptback-Wing Fighter-Airplane Model. NACA I?M L56Flga, 1956.
16. Hastings, Earl C., Jr.: Minimum Drag of Four Versions of a Swept- Wing Fighter Airplane Obtained From Flight Tests of Rocket-Boosted Models at Mach Numbers From 0.81 to 1.71. NACA RM L56E25a, 1956.
17. Larson, Terry J., Stillwell, Wendell H., and Armistead, Katharine H.: Static-Pressure Error Calibrations for Nose-Boom Airspeed Installa- tions of 17 Airplanes. NACA RM H57A02, 1957.
c
TABLE 1 . . PERTINENT CHARACTERISTICS OF TEST AIRPLANE
Wing (notincludingleading-edgechordlextension):
Area. sq f t . . . . . . . . . . . . . . . . . . . . . . . . . 375
. Span. f t . . . . . . . . . . . . . . . . . . . . . . . . . . 35.67
A s p e c t r a t i o . . . . . . . . . . . . . . . . . . . . . . . . 3.4
T a p e r r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . 0.247
3 Sweepback of quarter-chord line. deg . . . . . . . . . . . . 42.0
?
I Dihedral. deg . . . . . . . . . . . . . . . . . . . . . . . . -5.0
.
Geometric w i n g incidence. relative to fuselage reference line:
Cruise and high speed. deg . . . . . . . . . . . . . . . . -1.0
Take-offandlanding.deg . . . . . . . . . . . . . . . . .
7.0
Wing-hinge-point location. percent mean geometric chord . . . 39.58
Mean geometric chord. in . . . . . . . . . . . . . . . . . . . 141.4
A i r f o i l s e c t i o n p a r a l l e l t o p l a n e of symmetry:
Wing root . . . . . . . . . . . . . . . . . . . . . . NACA 65~006
Wing t i p . . . . . . . . . . . . . . . . . . . . . . NACA 65AoO5
Deflections of leading-edgedroop: Inboardsection:
Landingandtake.off. deg . . . . . . . . . . . . . . . . 25
Cruise. deg . . . . . . . . . . . . . . . . . . . . . . . 6.75
. . . . . . . . . . . . . . . . . . . . . 0
Highspeed. deg Outboard section:
Landingandtake.off. deg . . . . . . . . . . . . . . . . 27
Cruise. deg
. . . . . . . . . . . . . . . . . . . . . . . 7.0
Highspeed. deg . . . . . . . . . . . . . . . . . . . . . 0
Chord-extension area (both sides). sq f t . . . . . . . . . . 10.33
Center-section inboard flaps:
Area (both sides). sq f t . . . . . . . . . . . . . . . . . 13.44
Deflection for landing and take.off. deg . . . . . . . . . 20.0
Deflection for cruise and high speed. deg . . . . . . . . . 0
Ailerons : Chord. percentof wing chord:
Outboard . . . . . . . . . . . . . . . . . . . . . . . . 28.0
Inboard . . . . . . . . . . . . . . . . . . . . . . . . . -23.5
k e a . sq f t . . . . . . . . . . . . . . . . . . . . . . . . 20.78
‘Deflect ions : . .
”
High speed and cruise. deg . . . . . . . . . . . . . . . k13
Take-offandlanding:
Both ailerons drooped as flaps. deg . . . . . . . . . . 20
A6 ailerons. deg . . . . . . . . . . . . . . . . . . . +45-15
TABLE I.- PERTINENT CHARACTERISTICS OF TEST AIRPLANE . Concluded
Vertical stabilizer (based on area extending to horizontal tail center line. not including dorsal): Area. sq ft . . . . . . . . . . . . . . . . . . . . . . . . .
Span. ft . . . . . . . . . . . . . . . . . . . . . . . . . . 12.75
Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . .
1.5
Sweepbackofquarter-chordline.deg . . . . . . . . . . . . 45.0
Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . 0.25
Mean geometric chord. in . . . . . . . . . . . . . . . . . . . 114.8
Tail length. from 28 yo wing mean geometric chord to
25 y o vertical-tail mean geometric chord. in . . . . . . . . 168.9
Airfoil:
Waterline . . . . . . . . . . . . . . . . . Modified NACA 65A005.3
Tip . . . . . . . . . . . . . . . . . . . . Modified NACA 65AOO4
Rudder :
Area. sq ft . . . . . . . . . . . . . . . . . . . . . . . . . 12.56
Chord. constant. in . . . . . . . . . . . . . . . . . . . . . 21.28
Maximum deflections:
High speed and cruise. deg . . . . . . . . . . . . . . . . k6.0
Take-off and landing. deg . . . . . . . . . . . . . . . . . k17.0
Horizontal stabilizer (based on area extending to fuselage center line):
Area. sq ft . . . . . . . . . . . . . . . . . . . . . . . . . 93.4
Span.ft . . . . . . . . . . . . . . . . . . . . . . . . . . 18.1
Aspect rat io . . . . . . . . . . . . . . . . . . . . . . . . 3.5
Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . 0.148
Sweepback of quarter-chord line. deg . . . . . . . . . .
. . 45
Geometric dihedral. deg . . . . . . . . . . . . . . . . .
. . 5.417
Mean geometric chord. in . . . . . . . . . . . . . . . . .
. . 73.4
Tail length. from 28% wing mean geometric chord to
2 5 y 0 horizontal-tail mean geometric chord. in . . . . . . . 200.6
Maximum deflections:
Trailing edge down. deg . . . . . . . . . . . . . . . . . . 8
Trailing edge up. deg . . . . . . . . . . . . . . . . .
. . 32
Airfoil:
Root . . . . . . . . . . . . . . . . . . . . . . . . NACA 65~006
Tip . . . . . . . . . . . . . . . . . . . . . . . . . NACA 65A004
1 I U (a) Three-quarter front view. L-57-2099 Figure 1.- Testairplaneinthetake-off and landingconfiguration.
(b) Rear view. L-57-2102 Figure 1.- Concluded.
I 1 J L Figure 2.- Three-view drawing of the test airplane. All dimensions a r e i n inches.
;O 20 10 0 10 7.0 30 40 Left Right Lateral sticr: position, deg (a) Aileroncontrolsystem,cleancondition.
Figure 3 . - Characteristics of the aileron and ruddercontrolsystems as measured on t h e ground. L a t e r a l s t i c k l e n g t h = 18- inches.
. " "
""
. - I
: : : I
X 4c 10 0 10 20 30 40 Right Lateral stick wsition, deg (b) Aileron control system, landing condition.
Figure 3 . - Continued.
r
- - . - I
Landing condition I . ..
Clean condition
-
4 0 4 8 12 16 Left Right Rudder position, dee; (c) Rudder control system.
Figure 3.- Concluded.
A f t Longibdinal stick position, deg Figure 4.- Characteristics of the longitudinal control system showing the variation of stick force and stabilizer deflection with stick displacement. Longitudinal stick length = 22 inches.
0 NASA Flight Data Point
- .c - Calculated Using
Ref. 17 Fig No. 17 Chance TTought data
Calibrated Mach Number Mc - Mi + AM
-
040 . 5 0 . 60 70 .80 .90 1 . 0 0 1 . 1 0
Indicated Mach number, 5.- Machnumber calibration used to correct indicated Mach number t o c a l i b r a t e d Mach num- Figure ber.Calibrationdoesnotexist between M i = 0.96 and M i = 1.00 (shown by short-dash l i n e ) .
I I - 2 . 3 . 6 . 7 .8 .9 1 . 0 Calibrated Mach number (a) 20,000 f e e t .
Figure 6.- Variation of stabilizer position and stick force with calibrated Mach number.
2 . I r
Calibrated Mach number (b) 35,000 feet.
Figure 6.- Concluded.
I I
Calibrated airspeed , knots
. .
Figure 7.- Variation of horizontal stabilizer position required for trim with calibrated air- speed in the landing condition.
1 2 3 4 5 Norm1 acceleration, g units (a) M, = 0.90; a l t i t u d e = 35,500 feet;cruisedroop down.
Figure 8.- Variation of l o n g i t u d i n a l s t i c k p o s i t i o n , s t i c k f o r c e , and horizontal stabilizer positionwith normal acceleration.Tests made i n windup t u r n s .
0 1 2 3 4 5 Normal. acceleration, g u n i t s (b) Mc = 1.35; a l t i t u d e = 29,700 feet;cruise droop up.
Figure 8.- Concluded.
I
0 1 2 3 & 1,:ormdl acceleration, g u n i t s Figure 9.- Variation of longitudinal stick position, stick force, and horizontal stabilizer position with normal acceleration. Tests made in windup turns with cruise droopup at a Mach number of 0.90 and and altitude of 14,400 feet.
a 4 .6 .8 1 . 0 l a 2 1 . 6
Calibrated Phch number Figure 10.- Summaxy plot of the variation of horizontal-tail deflection per g as a function of Mach number f o r several altitudes and the two cruise-droop positions.
4c Full 2c Stick force, Lb A f t -10 C Stick position, de6
0 . L 10
T.E up IIorizontal stabilizer anzle, deg Normal acceleration, g 1 A n g u l a r velocity, rad/sec A n g u l a r acceleration, rad/sec2 -1.c 1 2 Time, 6ec (a) M , = 0.98; a l t i t u d e = 34,900 f e e t .
Figure 11.- Time h i s t o r i e s of rapid pull-up andholdmaneuvers.
I
43.
30 ~ pull S t i c k f o r c e , l b S t i c k position, de& 0 I /- .
cc I
Aft - : E
rl T.E. u p Horizontal stabilizer angle, deg
3 't
Normal acceleration, g Angular acceleration, rad/sec2 J 0 1 2 3 Time, sec .
(b) Mc = 0.93; a l t i t u d e = 13,760 f e e t .
Figure 11.- Concluded.
1 2
-
0 .h .8 1 . 2 1 . 6 2 .o Normal acceleration, g units (a) Vc = 140 knots; a l t i t u d e = 5,000 f e e t .
Figure 12.- V a r i a t i o n o f l o n g i t u d i n a l s t i c k - p o s i t i o n , s t i c k f o r c e , and h o r i z o n t a l s t a b i l i z e r p o s i t i o n w i t h normal a c c e l e r a t i o n i n t h e landing condition.
0 .4 .8 1 . 2 1 . 6 2 . 0 Normal acceleration, g u n i t s (b) V , = 197 knots;altitude = 11,000 f e e t .
Figure 12. - Concluded.
Damper Walking beam Figure 13.- Schematic drawing of the longitudinal control system of the test airplane. The enlarged drawings indicate the areas where measurements were made. An asterisk on the link- age indicates the point at which the measuring instrument was attached.
t % = 1.208 q i = 746 VI = 443.5 Ha = 34,300 M , = 1.134 q, = 547 V I = 3 8 5 HpL = 37,325 M, F 1.C qi = 423 vi = 341.5 Hpi = 34,400 M, = 0.953 q , = 387 vi = 280 ti - :lb,9W Pi - M, = 0.342 qi = 280 vi = 281 Hpi = 35.1% llnrmal accelerallon, g (a) Altitude = 35,000 f e e t .
Figure 14.- Variation of longitudinal stick position, walking beam l i n k - age position, structural feedback linkage position, and r i g h t h o r i - z o n t a l s t a b i l i z e r p o s i t i o n w i t h normal acceleration.Tests made i n windup t u r n s for various Mach numbers a t a l t i t u d e s of approximately 35,000 f e e t and 20,000 feet.
Stick """ Walking beam
"- Structural feedback
"" Right stabilizer Me = 0.990 Si = 788 Vi = 454.5 Hpi = 20,325 a, Mc = 0.936 tn Q = 719 rd " V i = 436 V Hpi = 19,475 Mc = 0.886 qi = 606 Vi = 403.5 Hpi = 20,580 1 2 3 4 5 6 Normal acceleration, g (b) Altitude = 20,000 feet.
Figure 14.- Concluded.
t P 49 .
cli
0 .842 .I398 3 23 3 67 d 547 co
L
1 2 3 4 5
Normal acceleration, g 15.- Summary plot of the structural deformation effects on the longitudinal control system as shown by the variation of the deforma- tion effects in terms of stick angle with normal acceleration as determined by the relative motion between the right horizontal sta- bilizer and the longitudinal stick position.
.7 . 9 .9 1 . 0 1 . 1 1 . 2 1.3 1.h Figure 16.- Characteristics of the short-period longitudinal oscilla- t i o n s showing the variation of the period, time to damp to one-half amplitude, and damping ratio as a function of Mach number at an altitude of approximately 35,000 feet.
.
-. h
C .d I
-..3
E ?
V - 1 . 2 - 1 . 6 -2 -a -10 .7 .0 .9 1 . 0 1 . 1 1 . 2 1 . 3 l . L 1 . 5 - Mach Nunter
Figure 17. - Variation of C and ( 2 % + Cm with Mach number as
ma 9
determined from the period and damping data presented in figure 16.
Data obtained at an altitude ofapproximately 35,000 feet.
I .? 1 . 0 1 . 1 .9 1 . 2 1.3 1 . h 1 . 5 Mach Numb e r Figure 18.- Variation of CN with Mach number as determined from the flight test data. Data a obtained atan altitude of approximately 35,000 feet.
n
=. 1
!iiL
-. 2
-. 3
I
-. u
.8 .9 1 . h .7 1 . 0 1 . 1 1 . 2 1.3 Mach Numb e r Figure 19.- Variation of aerodynamic-center location with Mach number as obtained from the flight test data. Data obtained at an altitude of approximately 35,000 feet.
ul w Rudder Stabilizer Ai-leron (total) 2 0 2 0 8 4 0 4 e Left Sideslip angle, deg R i g h t (a) Mc = 0.83; a l t i t u d e = 35,000 f e e t .
Figure 20.- Static directionaL stabili-ty and control characteristics of t h e t e s t a i r p l a n e as indicated by t h e v a r i a t i o n of control positions and forces with sideslip angle for various Mach numbers at a l t i t u d e s .
0 Rudder 0 S t a b i l i z e r Aileron ( t o t a l )
if
1 0 1 0 cr
J 20
8 4 0 4 8 Left Rig h t Sideslipangle, deg - ." . ~ (b) M , = 1.51; a l t i t u d e = 34,500 f e e t .
Figure 20.- Continued.
L
0 Rudder 0 Stabilizer , Obileron ( t o t a l )
s
M .A lz rl 0 0 k k a a 100 2 8 0 h 0 Left R i g h t Sideslipangle, deg (c) Mc = 0.83; altitude = 21,000 feet.
Figure 20.- Concluded.
lo-7-58~ -2 -4 e 8 .9 1 . 0 1 . 1 1 . 2 1.3
1 . 4 L.5 1 . 6
Calibrated Mach number 0 Rudder 0 Stabilizer O l i l e r o n (total .
a h 0 h 8 Right Left Sideslip angle, deg (a) V = 200 knots;altitude = 8,500 f e e t ; CN = 0.44; ai,f = -0.84'; ai ,w = 6.16'.
Figure 22.- 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 control characteristics of t h e t e s t a i r p l a n e i n t h e l a n d i n g c o n d i t i o n as indicated by the varia- - t i o n of control positions and forces with sideslip angle f o r three airspeeds.
-
0 Rudder 0 Stabilizer 0 Aileron ( t o t p l ) 8 h 0 h Left Sideslip angle, deg ( c ) V = 138 knots; a l t i t u d e =- 3,300 feet; CN = 0.99; ai,f = 10.1O; a i , w = 17.1O.
Figure 22.- Concluded.
I L Calibrated Mach number 1 . 2 .9 1 . 0 1 . 1 .b .6 . 7 1.3 Mach Number, M Figure 24.- Plot of the calculated and measured values of the time required to roll through a 90' angle of bank using maximum aileron deflection as a function of altitude and Mach numb Numerals designate time in seconds to roll through a 90' angle of bank.
I .
system .7 .9 .9 1 . 0 1 . 1 1 . 2 1 . 5 Mach Number Figure 25.- Characteristics of the lateral directional oscillations showing the period, time to damp to one-half amplitude,and damping ratio as a function of Mach number at an altitude of approximately 35,000 feet. The data are presented for the stabilization system both on - and off.
S t ; b z z a t i o n system
- - 0 off
n I
.? . r j .9 1 . 2 1 . 1 1 . 3 1.3 1 . 1 1
Mach iu’umber
Figure 26. - Roll-to-yaw r a t i o s as measured during the lateral direc-
t i o n a l o s c i l l a t i o n s f o r t h e s t a b i l i z a t i o n system both on and off.
Tests made at an a l t i t u d e of about 35,000 f e e t .
, 1 9P
-
2 . 5 2.b 2 . 0 1 . 6 II c1/2 1 . 2 .8 .h 0 . 2 .4 .6 .8 1 . 0
B
-
Ve
Figure 27. - Damping parameter - as a function of the parameter -. B
(3112 ve Data are presented for the stabilization systems on and off at an altitude of about 35,000 feet; also shown are the requirements as specified in reference 11.
.
cn cn Figure 28.- Variation of the static directional-stability parameter C , , with Mach number for the stabilization systems off at an altitude of 35,000 feet. Data " P obtained from the period and damping data presented in figure 25.
.
I 10- 7- 5 8 ~ .8 .9 1.0 1 . 1 1.2 1.3 1 . 4 Mach Nwher Figure 29.- Variation of the side-force parameter w i t h Mach number for the stabilization cyP systemoff at an a l t i t u d e ofabout 35,000 feet.
6 8 P u l l Stick force, l b 0 Fush 20 Aft Stickposition, deg F d T . E . up Horizontal stabilizer angle, deg Normal acceleration, g AnEular velocity, rad/sec 0 .a 0 2 6 10 17 Time, s e c (a) Attemptby t h e p i l o t t o trim the airplane rapidly.
Figure 3 0 . - Time h i s t o r y of attempt t o t r i m the airplane with the normal longitudinal t r i m system.
Stick.force, lb Push 20 .04 Angular velocity, r a u s e c - .04 0 2 4 6 8 10 Time, sec (b) Typical example of t h e pilot's trimming procedure.
Figure 30.- Concluded.
L e f t Aileron stick position,deg Right Aileron angle, deg T.E. down Compression Wing incidence strut loading, lb Tension 2000 L.E. up Wing incidence 3.5 angle,deg T.E. up Horizontal stabilizer angle,deg Normal acceleration,
I L I I I L 1 I
0 2 4 6 Time, sec Figure 31.- Time history of a successful operation of the wing from the landing condition to the clean condition. Tests made with the cruise
droop down at an indicated airspeed of 201 knots and an altitude of -
11,500 feet.
Left Aileron stick position, deg n Right Aileron angle, deg T.E. down
20 -
Compression 2000 - Wing incidence strut loading, l b 0 Tension 2000 - T.E. up Horizontal stabilizer angle, deg
5 1
Normal acceleration, g 0 2 4 6 Time, sec Mine incidence strut loading, lb 0 \ L.E. up W i n g incidence angle, deg 3.5 0 -
T . E . up 10 -
4-
I I I I I I I I - 2 4 6 8 10 1 2 Time, sec Figure 33.- Time h i s t o r y of the operation of t h e wing from t h e f u l l up tothecleanconditionduring a typicaltake-off. The indicated airspeed varies fromabout 190 knots to about 220 knots during the t e s t s .
10P
-
Left Aileron stick Right 5 I 4 a3 n Compression 2000
t"
I rl Wing incidence strut loading, lb Tension 2000
7 h
L.E. UP Wing incidence angle, deg 3.5
- \\
-
T.E. up 10 Horizontal stabilizer angle, deg
5 t
/ .
2 4 Time, sec Figure 34.- Time history of an unsuccessful attempt t o lower t h e wing t o thecleancondition. Tests m d e w i t h t h e c r u i s e droop up at an indicated airspeed of 197 knots at 11,500 feet a l t i t u d e .
Aileron stick . - - ~- - ~~
-
position, deg Right 10 - - Aileron angle, 0 l r L e f t c """""""""""~"""""" """" de9 - tRighi T.E. down lo - Compression 3000
Wing incidence - - . . ~. ~
strut loading, lb e -88 D m p DO= Tension 3000 - 3.6 - C r u i s s Dmop W L.E. up . - . . _ _ _ Wing incidence 0 angle, deg
-
T.E. UP 10 Horizontal stabilizer 'J angle, deg
t
Nonnal acceleration, g 1 -
/ __I Time. sec Figure 35.- Time h i s t o r y of the operation of t h e c r u i s e droop at an indicated airspeed of about 178 knots and an altitude of 15,800 feet.
Left 10 Aileron stick /L/ position, deg 0 - m r h t 10
rLeR
Ailel-on angle, 0 3 deg - Right - T.E. down 20
-
compression 3000 Wing incidence .
/ Strut loading, l b - Tension 3000 - 3.5 L.E. up Wing incidence - angle,deg 3.5 I - T . E . up 10
-
Horizontal stabilizer 5 angle, deg I I I I I I I I I I I n
-
0 2 4 6 8 10 12 Time, sec Figure 36.- Time history of the operation of theleading-edge droop from the full-up position to the landing-droop position with the wing i n t h e down and lockedposition.Testsconducted a t an indi- cated airspeed of about 1 8 6 knots a t an a l t i t u d e of l5,mO feet.
L - Aileron stick position. deg
Tension 3oco 1
4 6 8 10 32 0 2 34 36 38 Time. sec (a) Wing-up cycle; (b) Wing-down cycle; V i = 198 knots; 15,500 feet. vi = 181 knots; 16,000 feet.
Figure 37.- Time history of the operation of the variable-incidence w i n g with the leading-edge droop locked in the landing-droop position during the entire cycle.
NASA - Langley Field. Va.