Document
N A S A T E C H N I C A L N O T E
FLIGHT INVESTIGATION OF STABILITY
A N D CONTROL CHARACTERISTICS OF A
1 / 9-SCALE MODEL OF A FOUR-PROPELLER
TILT-WING V/STOL TRANSPORT
by WiZhhm A. Newsom, Jr., and Robert H. K i r b y
LangZey Research Center
LangZey Station, Hampton, Vu.
N A T I O N A L AERONAUTICS AND SPACE A D M I N I S T R A T I O N 0 WASHINGTON, D. C . 0 SEPTEMBER 1964 TECH LIBRARY KAFB. NM
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0079573 FLIGHT INVESTIGATION OF STABILITY AND CONTROL CHARACTERISTICS OF A 1/9-SCALE MODEL O F A FOUR-PROPELLER TILT-WING V/STOL TRANSPORT By William A. Newsom, Jr., and Robert H. Kirby Langley Research Center Langley Station, Hampton, Va.
Technical Film Supplement L-835 available on request.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For s a l e by tho O f f i c e of T e c h n i c a l Services, Department of Commorco,
Washington, D.C. 20230 -- Prico $1.25
mGHT INVESTIGATION O F STABILITY AND C O N T R O L CHARACTERISTICS O F A l/g-SCAL;E MODEL OF A FOUR-PROPEUXR TILT-WING V/STOL TRANSPORT By W i l l i a m A. Newsom, Jr., and Robert H. Kirby SUMMARY A f l i g h t investigation has been made t o study t h e s t a b i l i t y and control characteristics of a l/g-scale model of a four-propeller tilt-wing V/STOL trans- port airplane. The tests included hovering f l i g h t s i n and out of ground e f f e c t and l e v e l f l i g h t and descent conditions i n t h e t r a n s i t i o n speed range. N o arti- f i c i a l s t a b i l i z a t i o n w a s used i n any of t h e t e s t s . Even though the model w a s it s t a t i c a l l y and dynamically unstable f o r many of t h e f l i g h t - t e s t conditions, could generally be controlled and maneuvered easily. The descent t e s t s showed t h a t t h e configuration had at l e a s t a 6 O descent capability with no adverse e f f e c t s , and t h a t an additional 3 ' or 4' of descent angle w a s available before completely unacceptable flying q u a l i t i e s were encountered as a r e s u l t of wing s t a l l i n g . I n a l l f l i g h t regions, t h e minimum t o t a l control powers found t o be s a t i s f a c t o r y i n t h e model f l i g h t tests w e r e l e s s than t h e control powers planned f o r t h e full-scale a i r c r a f t .
INTRODUCTION An investigation t o study the low-speed dynamic s t a b i l i t y and control char- a c t e r i s t i c s of a four-propeller tilt-wing V/STOL transport airplane has been made at t h e NASA Langley Research Center using a l/g-scale model.
The wing i s pro- vided with a full-span double s l o t t e d f l a p which i s programed t o deflect as t h e wing incidence changes.
The investigation included free-flight tests i n s t i l l a i r f o r study of the vertical-take-off-and-landing and hovering-flight conditions and free-flight tests i n t h e Langley full-scale tunnel f o r study of slow constant-altitude t r a n s i t i o n s and simulated descending-flight conditions at low t r a n s i t i o n speeds The r e s u l t s were mainly q u a l i t a t i v e and consisted of p i l o t s ' observations and opinions of t h e behavior of t h e model.
SYMBOLS b wing span, f t Cn yawing-moment coefficient, Mz/qSb l o c a l wing chord, f t wing mean aerodynamic chord, f t model propeller diameter, f t height of model fuselage above ground ( e = Oo) moment of i n e r t i a about X body axis, slug-ft2 moment of i n e r t i a about Y body axis, slug-ft2 moment of i n e r t i a about Z body axis, slug-ft2 wing incidence, deg radius of gyration about X body axis, f t radius of gyration about Y body axis, f t radius of gyration about Z body axis, f t lift, l b l i f t i n hover out of ground e f f e c t r o l l i n g moment due t o r o l l angle, ft-lb/deg pitching moment due t o fuselage pitch angle, ft-lb/deg yawing moment, ft-lb yawing moment out of ground effect, f t - l b rate of r o l l , radians/sec dynamic pressure, l b / f t 2 wing area, f t 2 velocity, ft/sec weight, l b coordinate axes angle of attack of fuselage, deg angle of sideslip, deg 7 flight-path angle, deg E a aileron deflection, deg t o t a l f l a p angle, measured between wing chord and second element of 6f flap, deg 9 fuselage pitch angle, deg r o l l angle, deg APPARATUS AND TESTS Model General description.- Photographs of the 1/9-scale model used i n the inves- t i g a t i o n are presented as figure 1. Drawings of the model showing some of the more important dimensions are presented i n figure 2. The geometric characteris- t i c s of the model are l i s t e d i n table I. The variation of center of gravity with wing incidence f o r the model and f o r the airplane is shown i n figure 3 . The moments of i n e r t i a of the configuration were essentially constant throughout the wing incidence range, and the average values f o r the model (scaled up) are com- pared with those of the full-scale airplane i n table 11.
The four main propellers of the model were interconnected by a system of shafts and gear boxes and were driven by a pneumatic motor. The t a i l rotor w a s The wing w a s pivoted at the 30-percent driven by a separate pneumatic motor.
mean aerodynamic chord s t a t i o n and could be rotated by an e l e c t r i c motor between angles of incidence of 0 ' and 90° during f l i g h t . The wing was equipped with a slat along t h a t part of the leading edge t h a t w a s behind the up-going propeller The wing w a s also equipped with the 47-percent-chord double s l o t t e d blades.
f l a p shown i n figure 2(b) which w a s programed with a simple cam and follower t o deflect as the wing incidence changed. The programed variation o f f l a p deflec- t i o n with wing incidence is shown i n figure 4.
Control system f o r hovering f l i g h t . - In hovering, r o l l control w a s provided by d i f f e r e n t i a l l y changing the t o t a l blade pitch of the four main propellers, and yaw control was provided by d i f f e r e n t i a l l y deflecting the conventional ailerons at t h i s 90° wing angle. The ailerons were b u i l t into t h e rear element of t h e f l a p as shown i n figure 2(b) and were located on the two outboard segments of the f l a p as shown by the shaded area i n figure 2(a). Pitch trim was obtained by t o t a l blade pitch of the t a i l rotor and pitch control f o r maneuvering was pro- vided by a j e t mounted at the rear of the model. It should be pointed out t h a t on the airplane both pitch control and trim are obtained from the t a i l rotor, but on the model, f o r mechanical reasons, it w a s not desirable t o obtain control from the t a i l rotor. The controls were deflected by flicker-type ( f u l l on o r o f f ) trim control of the t a i l rotor which w a s pneumatic actuators except f o r the pitch actuated by an e l e c t r i c motor. The main propeller-blade pitch actuators were equipped with an integrating-type trjmmer t h a t trimmed the control a small amount each time the f l i c k e r control w a s given. The aileron actuators were mounted, f o r trim, on movable platforms driven by a small e l e c t r i c motor. The j e t reaction not equipped with a trimmer.
control used f o r p i t c h control w a s Control system f o r conventional forward f l i g h t . - I n conventional forward f l i g h t where t h e wing and propellers were at a tilt angle of Oo, the model had conventional ailerons and rudder f o r r o l l and yaw control. The rudder, however, did not provide s u f f i c i e n t yawing moment by itself; therefore, t h e yaw control w a s augmented i n t h e model tests by t h e use of d i f f e r e n t i a l blade p i t c h changes on the four main propellers. The j e t reaction control used f o r pitch control i n hovering w a s a l s o used throughout t h e investigation from hovering t o conventional forward f l i g h t . The model did have an all-movable horizontal t a i l t h a t w a s pro- gramed t o move as the wing incidence changed but t h e t a i l w a s not controlled by t h e p i l o t . The programed variation of t h e horizontal-tail incidence with wing incidence is shown i n figure 4.
Control system f o r t r a n s i t i o n f l i g h t . - I n t h e t r a n s i t i o n range t h e ailerons as t h e and t h e d i f f e r e n t i a l propeller p i t c h control interchange t h e i r function wing incidence changes. On the full-scale airplane a control mixing device is desired response t o t h e p i l o t ' s control movements.
used t o give t h e I n general, the propeller blade p i t c h and aileron control are mixed according t o t h e wing incidence s o t h a t lateral s t i c k always results i n a r o l l control and pedal dis- placement gives a y a w control. No such mechanical control mixer w a s used i n t h i s model investigation, however. The model p i l o t s were able t o use various combinations and amounts of these controls by e l e c t r i c a l switching of the f l i c k e r mechanisms and by ground adjustment of t h e amount of control given by t h e f l i c k e r mechanisms. The control moments used during t h e d i f f e r e n t f l i g h t conditions are presented subsequently.
T e s t Techniques The basic test setup used i n t h e present tests was e s s e n t i a l l y t h e same as t h a t used f o r a l l flight tests i n t h e Langley full-scale tunnel and is illus- trated i n figure 5 . An additional operator (not shown i n f i g . 3 ) was located near t h e p i t c h p i l o t t o control t h e wing incidence i n some of t h e tests. The t h e control t r i m motors, and t h e electric-control power f o r t h e wing tilt motor, solenoids was supplied through w i r e s ; and t h e air f o r t h e pneumatic motors, t h e jet-reaction control, and the control actuators was supplied through p l a s t i c tubes. These w i r e s and tubes were suspended f r o m t h e t o p of t h e tunnel and were taped t o a safety cable (1/16-inch braided a i r c r a f t cable) from a point about 15 feet above t h e model down t o t h e model i t s e l f . The safety cable, which was attached t o t h e fuselage near t h e model center of gravity, was used t o pre- vent crashes i n t h e event of a power o r control failure o r i n t h e event t h a t t h e p i l o t s l o s t control of t h e model. Separate p i l o t s are used t o control t h e model The reasons f o r using t h i s model f l i g h t technique i n i n pitch, r o l l , and yaw.
which the p i l o t i n g duties are divided i n preference t o t h e conventional single- p i l o t technique is explained i n d e t a i l i n reference 1. In forward (and descending) f l i g h t two p i l o t s are sometimes used, one p i l o t controlling both r o l l and yaw.
Tests t o study t h e level-flight t r a n s i t i o n characteristics of a model can be made i n t h e Langley fuJ-l-scale tunnel e i t h e r by continually increasing o r decreasing the tunnel airspeed u n t i l the t r a n s i t i o n i s completed or by holding the tunnel airspeed constant a t intermediate speeds f o r more careful study of any s t a b i l i t y and control characteristics or problems t h a t may be encountered.
It has been found i n previous work with tilt-wing V/STOL a i r c r a f t (see ref. 2 ) t h a t one of the most c r i t i c a l f l i g h t conditions i s the p a r t i a l l y tran- sitioned descent condition which w i l l probably be used for most landing approaches. I n order t h k t this condition might be studied i n the present investigation, t h e free-flight t e s t i n g technique i n the Langley full-scale tunnel has been extended t o permit tests representing the descent condition t o be made i n the horizontal airstream of the tunnel. The factors involved i n the simulation of a descent condition a r e i l l u s t r a t e d i n figure 6. This figure shows t h e balance of forces involved i n actual descent a t the l e f t and i n the simulated descent a t the right. For the actual descent case, the l i f t , the drag being balanced by the forward drag, and weight forces a r e i n balance, component of the weight acting along the f l i g h t path. For the simulated descent condition i n the horizontal airstream of the tunnel, t h e model i s flown with effectively the same l i f t and drag, but the drag cannot be balanced by a component of the weight and must be balanced by some thrust force t h a t i s inde- pendent of the normal airplane l i f t and propulsion system. A small high-pressurc compressed-air j e t exhausting from the rear of the model where the aerodynamic interference effects would be negligible was used. I n t h i s way the aerodynamic effects of descending or decelerating f l i g h t , which a r e very important f o r many V/STOL a i r c r a f t types, can be simulated with the model i n l e v e l f l i g h t i n the tunnel. This method of simulation, however, does not account f o r the effects of descent angle on c l a s s i c dynamic l a t e r a l s t a b i l i t y , but fortunately these effects a r e small for the descent angles l i k e l y t o be encountered i n normal operation and a r e of much l e s s importance than the aerodynamic effects which can be correctly simulated.
For hovering t e s t s , a t e s t setup very similar t o t h a t shown i n figure 5 i s made i n a special hovering t e s t area located i n a large enclosure where the p i l o t s can be stationed closer t o t h e model than i s possible i n the t e s t section.
It has been found very desirable, particularly during t e s t s i n which the model i s flown very close t o the ground, f o r t h e p i l o t s t o be near the model so t h a t they can notice more readily and correct f o r s l i g h t changes i n model a t t i t u d e and altitude.
Tests The free-flight investigation included tests a t three different f l i g h t conditions: (1) hovering (both i n and out of ground e f f e c t ) , (2) steady l e v e l forward f l i g h t a t a = 0 ' (over t h e whole t r a n s i t i o n range from hovering t o iw = O o ) , and (3) simulated descent f l i g h t (at iw = 20°, 30°, 40°, and 5 0 ' f o r descent angles of Oo, 5 O , 'f', loo, l3O, and l5O).
The s t a b i l i t y , control- l a b i l i t y , and the general f l i g h t behavior were determined qualitatively from the p i l o t s ' observations; and-motion-picture records of t h e f l i g h t t e s t s were made as an a i d i n the p i l o t s ' evaluation and t o supply some quantitative data on the modelmotions.
N o a r t i f i c i a l s t a b i l i z a t i o n was used i n any of the t e s t s . The basic sta- b i l i t y of the model was studied, i n each f l i g h t condition, by having two of the p i l o t s controlling t h e model as steadily as possible ( a f t e r a trimmed condition had been established) while t h e t h i r d p i l o t made the t e s t s required t o determine the s t a b i l i t y of a particular phase of the model motion. I n t h a t manner, f o r example, the stick-fixed pitching o r rolling motions of the model were deter- mined. The controllability was determined i n the same manner by each p i l o t i n turn varying h i s control power t o determine the amount of control required f o r steady flying and f o r performing various maneuvers. The basic s t a b i l i t y or con- t r o l characteristics of a model do not, however, give the complete picture of the model f l i g h t characteristics; therefore, t h e model p i l o t s also assessed i t s general f l i g h t behavior, including the e f f e c t s of such factors as wing s t a l l i n g .
A few force tests were made, i n addition t o the free-flight t e s t s , t o help document some of t h e aerodynamic and s t a b i l i t y and control characteristics of the model.
RESULTS AND DISCUSSION A motion-picture film supplement ( ~ 8 3 5 ) t o t h i s report has been prepared and i s available on loan.
A request card form and a description of the film w i l l be found a t the back of t h i s document.
I n reviewing t h e r e s u l t s of the f l i g h t tests, it should be remembered t h a t , ~ a s shown i n t a b l e 11, t h e scaled-up weight and i n e r t i a characteristics of the test model were high i n comparison with t h e full-scale values. The radii of gyration of the model however were of approximately the right magnitude. These high m a s s characteristics of the model could have affected the detailed r e s u l t s of this investigation; f o r example, they could cause s l i g h t changes i n the period of the hovering oscillations o r changes i n t h e damping of the l a t e r a l oscillatory motions i n forward f l i g h t . It i s f e l t , however, that since the periods of t h e motions experienced with t h i s model were relatively long, the conclusions reached from the f l i g h t t e s t s are valid and were not appreciably affected by the high mass characteristics. The r e s u l t s of the forward-flight t e s t s would, however, apply directly t o f l i g h t a t an a l t i t u d e of about 11,000 f e e t because of t h e relationship between correct and actual wing loadings.
A l l the r e s u l t s a r e f o r the case of t h e a i r c r a f t without a r t i f i c i a l sta- b i l i z a t i o n since no a r t i f i c i a l s t a b i l i z a t i o n w a s used a t any time during the t e s t s .
Hovering Out of Ground Effect The f l i g h t t e s t s i n still air out of ground effect t o determine the basic s t a b i l i t y i n hovering f l i g h t showed that, as expected with a tilt-wing configu- ration, the model had unstable control-fixed oscillations i n pitch and r o l l and w a s neutrally stable i n yaw.
Examples of t h e motions encountered i n pitch and r o l l a r e shown by t h e t i m e h i s t o r i e s presented i n figures 7 and 8. These time h i s t o r i e s were obtained from motion-picture records of the model f l i g h t s . The pitching o s c i l l a t i o n was a predominantly angular motion without much translation, whereas the rolling o s c i l l a t i o n had a substantial translation accompanying t h e angular motion. The period of t h e pitching o s c i l l a t i o n was about 3.4 seconds These values and the period of the r o l l i n g o s c i l l a t i o n was about 6 seconds.
scale up t o about 10 and 18 seconds, respectively, f o r the full-scale airplane.
In s p i t e of the f a c t t h a t t h e model had unstable control-fixed pitching and rolling oscillations i n hovering f l i g h t out of ground effect, the p i l o t s f e l t t h a t the general f l i g h t behavior of the model w a s good. The model could be flown smoothly and could be maneuvered readily from one position t o another.
One reason t h a t t h e model was easy t o control i n s p i t e of the unstable oscilla- tions w a s t h a t t h e motions were relatively slow i n s t a r t i n g and were not easily excited by outside e f f e c t s such as gust disturbances o r movements of the control and power cable. Another reason t h a t t h e model was easy t o control was t h a t the period of the o s c i l l a t i o n was very long and thus the p i l o t w a s not conscious of i t s presence i n normal flying. This same general type o f r e s u l t was obtained a t both model scale and f u l l scale with the VZ-2 research airplane as indicated by references 3 and 4.
I n t h e f l i g h t tests t o determine how much control power w a s required f o r steady f l i g h t and f o r performing various maneuvers, t h e model p i l o t s found that l e s s control acceleration was required f o r satisfactory controllability than i s provided on the full-scale a i r c r a f t . The full-scale a i r c r a f t hovering con- trols should provide accelerations of about 0.80 radian/sec2 i n pitch, Actually, the model 1.08 radians/sec2 i n r o l l and 0.53 radian/sec2 i n yaw.
p i l o t s found t h a t 60 percent of t h e scaled-down value i n pitch, 50 percent of the scaled-down value i n r o l l , and 40 percent of t h e scaled-down value i n yaw were adequate f o r performing any t e s t maneuver required of t h e model. It has been found, as pointed out i n reference 1, that flying model r e s u l t s generally correlate well with full-scale f l i g h t t e s t results on t h e control power required i n pitch and r o l l , but the yaw-control requirements have not shown correlation with full-scale experience. The yaw-control task i n model flying i s mainly one of simple alinement under steady flying conditions and does not involve gusts, operation i n cross winds, maneuvering i n yaw, o r other disturbances and require- ments found i n f i l l - s c a l e t e s t s .
Hovering i n Ground Effect I n addition t o t h e hovering f l i g h t t e s t s made out of ground effect, a number of f l i g h t s were made t o study t h e effect of close proximity t o the ground on the model characteristics. These f l i g h t t e s t s showed t h a t near the ground the model w a s somewhat easier t o fly i n r o l l and pitch than it w a s out of ground effect.
The unstable control-fixed pitching motion that w a s present a t a l t i t u d e seemed t o become stable at very low heights when the wheels were about t o touch the ground. This characteristic is indicated by figure 9 which shows a t i m e history of t h e stick-fixed pitching motions of t h e model when hovering near the ground.
This figure shows t h e stick-fixed motion t o be a somewhat random motion of small amplitude when the wheels were almost touching the ground during the first part of the f l i g h t . When the model rose t o a s l i g h t l y greater height above the ground a f t e r about 7 seconds, the motion developed into a fixed amplitude oscillation.
After about 18 seconds of f l i g h t , t h e motion damped when the wheels touched t h e ground but built up again t o a larger amplitude motion. The effect of ground proximity on r o l l was l e s s pronounced than t h a t on pitch. The model p i l o t could not detect any appreciable change i n s t a b i l i t y but f e l t t h a t the r o l l i n g phase of the model motion became s l i g h t l y easier t o control a s t h e model neared the ground. The variation of s t a t i c s t a b i l i t y with height above the ground, as measured i n force t e s t s , i s shown i n figure 10. These data show t h a t the model had a s l i g h t amount of s t a t i c s t a b i l i t y i n p i t c h and r o l l as the model approached the ground; this condition probably accounts f o r t h e improved dynamic s t a b i l i t y and controllability.
Unlike the r o l l i n g and pitching motions of the model, t h e yawing motions became somewhat more d i f f i c u l t t o control as t h e model neared t h e ground. The model experienced e r r a t i c yaw disturbances which were apparently caused by t h e e r r a t i c nature of the recirculating slipstream which was aggravated by the other modelmotions. Although not large i n magnitude, these disturbances resulted i n greater p i l o t e f f o r t being required t o hold a desired heading f o r hovering near the ground. The yaw p i l o t also noticed a reduction i n the yawing moment produced by the ailerons near the ground, but he did not f e e l t h a t this loss of effective- ness w a s the major factor i n the increased control e f f o r t required. For most of the f l i g h t s t h e control used w a s t h e same S O o deflection t h a t was used out of ground effect, but a few f l i g h t t e s t s were made with a yaw control deflection of k4Oo. The increased control power gave a more positive yaw control and enabled the p i l o t t o correct quickly f o r the e r r a t i c disturbances but did not materially reduce the p i l o t e f f o r t o r concentration required t o hold a yaw heading. Fig- ure 1 1 shows the loss of effectiveness of t h e ailerons i n ground proximity f o r the model as measured i n force t e s t s . These data show t h a t the yaw control effectiveness of the ailerons w a s only about one-half as great when the wheels were almost touching the ground as when the model was out of ground effect.
Take-off and landing f l i g h t tests showed no apparent changes i n t r i m with a l t i t u d e about any of t h e axes. With the controls perfectly trimmed f o r out of ground effect, hovering several t e s t s were made which showed no tendency of the model t o move e i t h e r forward o r backward a t take-off.
A very d e f i n i t e ground effect on the model l i f t was noted i n the landing t e s t s . If the model thrust w a s reduced slightly so t h a t a slow v e r t i c a l descent was s t a r t e d from hovering f l i g h t , the model would descend down t o a certain point and would descend no f a r t h e r u n t i l the thrust w a s f u r t h e r reduced. If the descent w a s made at a s l i g h t l y f a s t e r r a t e , t h e model would rebound s l i g h t l y as if it were bouncing on a spring. If the descent r a t e was too f a s t , however, the on down i n s p i t e of t h e favorable ground effect momentum would carry t h e model Figure 12 shows the variation of l i f t with and it would s t r i k e the ground.
ground proximity obtained from force t e s t s of t h i s model. These data show a 20-percent increase i n model l i f t with constant propeller speed a t a value of of 0.25, which i s approximately the height a t which t h e wheels would touch h/D down f o r the full-scale airplane with the shock struts f u l l y extended. Analysis of the data of reference 5 indicates that p r a c t i c a l l y none of t h i s increase i n l i f t due t o ground proximity i s caused by increase i n the propeller thrust; therefore, it can be presumed t h a t almost the e n t i r e 20 percent increase i n l i f t w a s caused by an upload on the bottom of the fuselage, t h e source of which i s explained i n references 5 and 6.
Level Flight i n Transition Longitudinal s t a b i l i t y . - The basic s t a b i l i t y of t h e model throughout t h e t r a n s i t i o n f l i g h t range w a s determined during constant airspeed f l i g h t t e s t s a = Oo. with t h e model trimmed f o r f l i g h t a t 3kamples of t h e type of motions experienced a r e shown i n f i g u r e 13 which presents t i m e h i s t o r i e s of the control- fixed pitching motions f o r wing incidence angles representing four different a i r - speeds. The curves show that, as noted previously, t h e control-fixed motion i n hovering w a s an unstable oscillation. A t a wing incidence of 6 5 O , l i t t l e differ- ence w a s noticed i n the motion since t h e model w a s at a very low forward speed because t h e programed f l a p w a s being deflected during t h i s wing-incidence change.
A t lower wing incidences, t h e motions became less unstable and t h e period of t h e o s c i l l a t i o n became very long. I n f a c t , t h e unstable motions at lower wing angles a t the p i l o t when he w a s f l y i n g t h e model i n t h e nor- were not noticeable a l l t o m a l manner. For instance, the o s c i l l a t i o n a t a wing incidence of 2 5 ' had a very long period (about 6 seconds model scale) and without looking carefully f o r t h e o s c i l l a t i o n at constant forward speed, t h e p i l o t would not ordinarily distinguish it from t h e normal gust, o r other disturbances that t h e model experiences i n f l i g h t tests. A t t h e lowest wing angle (iw = 1 0 ' ) t h e t i m e history of figure 13 seems t o show that t h e two o s c i l l a t o r y modes normal f o r conventional forward f l i g h t a r e beginning t o appear - t h e short-period o s c i l l a t i o n shows up i n t h e p i t c h angle record, and t h e long-period phugoid o s c i l l a t i o n seems t o be appearing i n t h e longitudinal and v e r t i c a l displacement traces. This progressive change from a longitudinally unstable t o an apparently stable f l i g h t condition as t h e t r a n s i t i o n progresses from hovering t o forward f l i g h t i s t y p i c a l of other tilt- wing configurations such as that of reference 7.
Lateral s t a b i l i t y . - I n the t r a n s i t i o n range, t h e model w a s even easier t o fly i n r o l l than it had been i n hovering. I n f a c t , as soon as t h e model s t a r t e d i n t o t r a n s i t i o n from hovering, t h e r o l l control became noticeably easier. This r e s u l t w a s evidently caused by t h e f a c t that t h e model w a s stable i n r o l l i n t h e t r a n s i t i o n range of f l i g h t instead of having an unstable o s c i l l a t i o n as it had i n hovering, and that t h e motions resulting from gusts o r control disturbances consequently damped out instead of exciting an unstable oscillation. These characteristics w e r e observed i n f l i g h t tests which w e r e made t o study t h e control-fixed l a t e r a l motions i n t h e t r a n s i t i o n range. I n these t e s t s , t h e model w a s trimmed as carefully as possible and then t h e r o l l and yaw p i l o t s stopped giving control so t h a t t h e controls remained fixed at t h i s trim setting.
A t a l l angles of wing incidence t e s t e d below iw = TO0, the resulting model motion w a s a slow sidewise divergence with l i t t l e yawing and no observable rolling. This type of motion might have been a s l i g h t aperiodic divergence o r might have been caused by some s m a l l remaining out-of-trim s e t t i n g of t h e yaw control, but there w a s c l e a r l y no significant degree of o s c i l l a t o r y i n s t a b i l i t y .
I n order t o investigate t h e o s c i l l a t o r y s t a b i l i t y characteristics further, some additional f l i g h t s were made a t angles of wing incidence from 800 t o 20° i n which, after t h e trimmed f l i g h t condition w a s established, t h e model w a s delib- erately disturbed by using t h e controls t o impart a combined r o l l i n g and yawing motion. Each t i m e , a f t e r t h e controls became fixed, t h e r o l l i n g and yawing motions damped out quickly but then t h e model performed the same sidewise trans- l a t i o n a l divergence (with some yawing) noted i n the previous control-fixed tests.
The f l i g h t t e s t s discussed, as w e l l a s normal controlled t r a n s i t i o n f l i g h t t e s t s , indicated t h a t the model had a region of neutral directional s t a b i l i t y f o r small angles of s i d e s l i p over the e n t i r e t r a n s i t i o n speed range. This sta- b i l i t y problem appeared i n controlled f l i g h t as a tendency of the model t o trim a t a small sideslip angle, i n e i t h e r right o r l e f t sideslips, which w a s objec- tionable t o t h e p i l o t s . Figure 14 presents t h e r e s u l t s of force t e s t s of t h e present model which show neutral directional s t a b i l i t y f o r a range of sideslip angles of about 6 O , but unpublished r e s u l t s from tests made with a l / l l - s c a l e model i n the Langley 7- by 10-foot tunnel indicated a s l i g h t amount of direc- t i o n a l s t a b i l i t y over the e n t i r e sideslip range and did not show a f l a t spot i n t h e directional s t a b i l i t y curve. This l / l l - s c a l e model, however, incorporated a number of minor changes i n the configuration t h a t w e r e made a f t e r the con- struction of the l/g-scale flying model had been completed. By temporarily modiQing the l/g-scale model, force t e s t s were made which showed that the f l a t spot i n the directional s t a b i l i t y curves could be eliminated with the flying model i f the gaps between the fuselage and the wing f l a p were sealed. (See f i g . 14.) The data of figure 14 also show t h a t the other modifications d i d not significantly change t h e directional s t a b i l i t y of the model. These other modi- fications consisted mainly of changes t o t h e f i l l e t a t t h e juncture of the ver- t i c a l t a i l , pitch fan support boom, and the fuselage.
I n order t o check t h e effect of sealing t h e f l a p gaps on the dynamic behavior of the model, f l i g h t t e s t s were made a t various angles of wing inci- dence with the gaps between t h e fuselage and t h e wing f l a p sealed. The f l i g h t characteristics of the model w e r e found i n these t e s t s t o be essentially unchanged from the unsealed condition and t h e directional s t a b i l i t y was s t i l l considered by the p i l o t t o be undesirably low. Because it was very awkward t o s e a l t h e flaps on t h e flying model and since very l i t t l e difference i n the flying characteristics resulted from the modification, the f l i g h t investigation w a s continued with t h e gap unsealed.
During the previously mentioned s e r i e s of force t e s t s made on the free- a few t e s t s were made with a larger v e r t i c a l t a i l . The area of f l i g h t model, the v e r t i c a l t a i l w a s increased 39 percent by an addition t o the leading edge and top of the f i n a s shown by the dashed l i n e s i n figure 2, and the t e s t s were made wlth the gaps between t h e f l a p and fuselage open. Figure 15 shows the e f f e c t of the larger v e r t i c a l t a i l on the directional s t a b i l i t y at iw = Oo, loo, and 20'. I n addition t o eliminating the neutral s t a b i l i t y a t small sideslip angles, the larger v e r t i c a l t a i l gave increased directional s t a b i l i t y over the e n t i r e range of s i d e s l i p angles. I n f l i g h t t e s t s made by using the large verti- the f l i g h t characteristics of t h e model were much improved by the c a l tail, increase i n directional s t a b i l i t y and there was no noticeable tendency on t h e p a r t of the model t o s i d e s l i p even at wing incidence angles as high as 5 0 ' where t h e airspeed w a s becoming f a i r l y low.
Descending Flight i n Transition Ekperience with the VZ-2 tilt-wing research a i r c r a f t , reported i n refer- ence 4, has shown that i n the reduced-power descending-flight conditions i n the a tendency t o s t a l l and t h a t t h i s s t a l l t r a n s i t i o n speed range, the wing has leads t o buffeting, abrupt wing dropping, and generally e r r a t i c , wallowing motions. These r e s u l t s of the s t a l l i n g were found t o become so severe that they effectively limited the r a t e of descent t h a t the p i l o t w a s willing t o use. This limitation can be very serious from an operational standpoint since it tends t o occur i n the speed range corresponding t o the landing-approach condition where high r a t e s of descent must be maintained t o take advantage of the short-field landing capability of V/STOL a i r c r a f t . Free-flight model t e s t s , reported i n reference 8, gave reasonably good agreement with the f'ull-scale f l i g h t t e s t s i n regard t o the wing s t a l l i n g and the limitations imposed on the operation of the a i r c r a f t by the wing-dropping and e r r a t i c , wallowing motions associated with the stalling. The buffeting, however, w a s not detected on t h e model which w a s not instrumented and was remotely controlled so that t h e p i l o t did not f e e l the buffeting. The free-flight model t e s t s therefore did not give the whole answer but seemed t o give t h e most important r e s u l t s with regard t o the seriousness of the wing s t a l l i n g . The characteristics of the present model were therefore studied very carefully with regard t o t h i s important problem.
The characteristics of t h e present basic model i n simulated descending f l i g h t were investigated over a wing incidence range of 20° t o TO0. The wing f l a p deflection grogramed as scheduled on the full-scale airplane (see f i g . 4) resulted i n a 60 f l a p deflection over most of the wing incidence range inves- tigated. A t each t e s t condition, the model was assigned a f l i g h t rating according t o t h e flying-model pilot-rating system shown i n table III. This model rating system i s shown and compared with the Cooper rating system since the intent of the model r a t i n g system i s t o consider the type of behavior of the model t h a t wodd represent insofar as possible the behavior required of an a i r - plane t o meet a l l the conditions given under the Cooper rating system. The ratings f o r the model are limited t o t h e s t a b i l i t y and control aspects of flying qualities since the remote-control p i l o t is unable t o sense the buffeting The p i l o t ratings obtained i n the t e s t s a r e shown i n figure 16 on a p l o t of f l i g h t path (or descent) angle against wing incidence. Ratings were obtained a t angles of wing incidence of 20°, 30°, 40°, and 50' f o r descent angles of Oo, 5O,
70, loo, l 3 ' , and 1 5 ' . The ratings shown i n figure 16 a r e overall ratings
obtained from t h e individual ratings on l a t e r a l , directional, longitudinal, and power characteristics (and f o r that reason the longitudinal and l a t e r a l s t a b i l i t y and control characteristics are not discussed separately as i n other sections of the report). A t each t e s t point, two ratings were obtained: (1) a rating of the behavior of the model when reasonably smooth and steady f i i g h t was maintained and (2) a rating f o r disturbed f l i g h t a f t e r the model had been intentionally given a large disturbance o r had been allowed t o build up i t s own large-amplitude dis- turbed motion. A t s m a l l descent angles, the model was very stable and had t o be disturbed intentionally with the controls, a f t e r which t h e disturbed motion danrped out quickly; therefore, f o r these conditions, there w a s no difference between the two ratings and only one rating i s shown i n figure 16. At the greatest descent angles, steady f l i g h t was not possible so only a disturbed- f l i g h t rating was given as indicated i n figure 16.
Figure 17 presents a summary of the p i l o t s ' opinions of the flying qualities of t h e model i n the form of boundaries obtained from the ratings of figure 16.
Above the dotted area i n figure 17, t h e model's characteristics were satisfac- tory and, i n f a c t , no difference from l e v e l f l i g h t was detected even when the model was intentionally disturbed. A s the descent angle w a s increased i n the
I
dotted area of figure 17, the model required more and more p i l o t attention t o the controls. A t the highest descent angles i n t h e dotted area, the l a t e r a l oscillations persisted f o r several cycles a f t e r a disturbance. I n those condi- tions intermittent s t a l l i n g of a p a r t of the wing could be observed from tufts on the wing. I n t h e hatched area of figure 17, the model experienced extensive wing s t a l l i n g which caused abrupt wing dropping, abrupt losses i n height, and t h e generally e r r a t i c , wallowing motions normally associated with wing s t a l l .
The model's flying q u a l i t i e s were unacceptable i n t h i s region.
condition Figures 18 and 19 a r e presented t o i l l u s t r a t e f o r the iw = 3 0 ° the types of f l i g h t characteristics encountered i n the descent t e s t s . Figure 18 shows t i m e h i s t o r i e s (from motion-picture records) of the l a t e r a l motions per- formed by the model while t h e p i l o t s were attempting t o make a smooth and steady I n l e v e l f l i g h t the model was controlled f l i g h t a t descent angles of 0 ' and 13'.
The e r r a t i c , very easy t o fly and required only occasional corrective control.
large-amplitude motions a t a descent angle of l 3 ' , however, were extremely d i f f i - c u l t t o control; and, i n f a c t , control of the model was l o s t a t times during the t e s t s . Figure 19 shows t i m e h i s t o r i e s obtained from f l i g h t s made a t 7 = - 5 O , - 7 O , and -loo t o study the motions performed by the model a f t e r it had been intentionally disturbed from a smooth flying condition by t h e p i l o t . A t 7 = - 5 O , three long control pulses were used by the p i l o t t o s e t up the motion, and the ensuing motion w a s so highly damped t h a t very l i t t l e p i l o t e f f o r t was needed t o
reestablish steady f l i g h t . 7 = -7, only two rapid control pulses were A t
needed t o s t a r t the motion but the motion w a s s t i l l mild enough so t h a t the p i l o t was able t o reestablish steady f l i g h t f a i r l y quickly. A t 7 = - 1 0 ' only one control pulse resulted i n the e r r a t i c , wallowing motions shown i n the figure.
These motions persisted i n s p i t e of t h e p i l o t ' s e f f o r t s t o reestablish steady flying conditions.
Several aspects of the behavior of the model do not show up i n t h e simple ratings. First, it should be noted i n figure 16 t h a t a rating of 4 was obtained This rating does not mean t h a t disturbances or f o r l e v e l f l i g h t a t iw = 50°.
wing s t a l l i n g were noticed i n t h i s condition but r e f l e c t s the f a c t t h a t a t the lower airspeeds the model did not have as much s t a b i l i t y a s a t the higher speeds and more p i l o t attention w a s required, A second point i s t h a t at times, during f l i g h t s at high-rate-of-descent conditions a t iw = 30°, t h e model would drop i n height abruptly without any appreciable e f f e c t on t h e l a t e r a l f l i g h t character- i s t i c s being noted. This abrupt loss i n height w a s a new type of motion not previously experienced i n t h e VZ-2 model t e s t s . Observation of the tufts on t h e wing showed that this abrupt dropping w a s caused by a sudden symmetrical s t a l l over a large p a r t of the wing. The last point t h a t should be brought out i n addition t o the simple ratings i s that a t high descent angles, somewhat dif- ferent model motions were obtained a t low angles of wing incidence than at high angles of wing incidence. iw = 20°, For example, as shown i n figure 16, a t steady f l i g h t could be achieved very easily and t h e tufts showed no apparent However, if a disturbance s t a l l i n g with descent angles as great a s 10'.
occurred at this point, the resulting abrupt wing dropping and generally e r r a t i c , wallowing motions of the model were very d i f f i c u l t t o control and a rating of 7 resulted. A t iw = 50°, however, there was not much difference between steady and intentionally disturbed f l i g h t at any descent angle. Although t h e tufts at descent angles as low as showed disturbed flow on the wing f o r steady f l i g h t 7 O , the model motions were not appreciably affected u n t i l the descent angle 1 1 ' . This effect might be expected since the high exceeded a value of about i w = 50' resulted incidence of the thrust l i n e and the high f l a p deflection a t i n most of the weight being supported by power rather than by wing l i f t so t h a t wing s t a l l affected only a very small part of the t o t a l l i f t .
A few t e s t s were made with the larger v e r t i c a l t a i l i n s t a l l e d on the model.
These t e s t s did not cover a l l the descent t e s t conditions t h a t were covered with the basic model but did cover enough conditions t o indicate t h a t the larger t a i l did not appreciably improve t h e behavior of the model a t the descent conditions i n which wing s t a l l i n g was causing the behavior t o be unsatisfactory.
I n summary, the model had at l e a s t a 6 ' descent capability with no adverse effects.
Another 3 ' o r bo of descent w a s available as a safety margin before completely unacceptable flying q u a l i t i e s were encountered. It should be pointed out again t h a t buffet e f f e c t s could not be evaluated i n these t e s t s . It might be inferred, however, t h a t since no disturbed flow could be detected on the wing at descent angles of 6 O o r l e s s , buffeting would not be expected t o cause any trouble i n t h i s f l i g h t region.
Evaluation of Control Power Required lrongitudinal control.- As mentioned previously, the pitch J e t w a s used throughout the f l i g h t range t o provide the longitudinal control required f o r maneuvering while t h e longitudinal t r i m required was provided by the t a i l rotor.
Figure 20 shows the longitudinal control power, i n excess of t h a t required f o r t r i m , planned f o r the airplane compared with the p i t c h J e t longitudinal control power (scaled up t o full-scale values) required on the model. The longitudinal control used on the model, which was l e s s than t h a t available f o r the full-scale airplane i n a l l cases, w a s found t o be adequate f o r any of the t e s t conditions including some rather abrupt maneuvering i n both l e v e l and descending f l i g h t .
Lateral control.- I n the transition-flight mode, the full-scale a i r c r a f t has a control mixing device which provides, a t each angle of wing incidence, a predetermined combination of propeller p i t c h and aileron deflection i n response t o a r o l l or yaw control from the p i l o t . The controls were not mechanically phased on the model but t h e r o l l and yaw p i l o t s could command preselected amounts and combinations of control moment during t h e t r a n s i t i o n i n order t o study the control requirements. Figure 2 1 shows t h e planned control powers f o r full l a t - eral s t i c k control and f u l l rudder pedal control on the full-scale a i r c r a f t , i n terms of angular accelerations, along with the control powers found t o be required during the present model t e s t s (includfng descending f l i g h t ) scaled up t o full-scale values. I n a l l cases the maximum control powers found desirable by the model p i l o t s were l e s s than t h e planned a i r c r a f t values. Also shown i n figure 2 1 are the helicopter control power requirements as s e t f o r t h i n the military specification of reference 9 and a point indicating the l a t e r a l con- t r o l power requfred f o r r o l l a t the higher forward speed.
The roll-control requirements determined with the model a r e i n good agree- ment with the helicopter requirements a t the low-speed end of the t r a n s i t i o n range and with t h e normal airplane requirements a t the high-speed end of the t r a n s i t i o n range. The yaw-control power required i n the model t e s t s , however, w a s much l e s s than the helicopter specification. It should be noted again t h a t yaw-control power required i n model t e s t i n g has not shown correlation with f u l l - scale experience f o r the hovering condition, evidently because the task i s simpler i n the model t e s t s . The model f l i g h t t e s t s did indicate, however, t h a t even though force t e s t s had shown t h a t the rudder w a s providing as much yawing- control moment as might be expected, the model could not be flown by using only the aileron and rudder as a coordinated control even at the highest speeds i n these t e s t s (iw = 0 ' with 6f = 30'). Analysis of the force-test data and the film records of the f l i g h t t e s t s indicated t h a t two factors were involved i n the apparent lack of rudder effectiveness. F i r s t , the adverse yawing moment caused by aileron deflection was apparently so large t h a t the yawing-control moment provided by the rudder w a s completely ineffective f o r opposing the adverse aileron yawing moment. Second, the film records showed t h a t when, f o r example, a right rudder control was given while the ailerons remained fixed, the large side force and adverse rolling moment due t o rudder deflection caused the model t o first r o l l and s l i d e t o the l e f t before it would f i n a l l y yaw and r o l l t o the right and start i n t o the intended right turn. Since the rudder w a s ineffective i n f l i g h t f o r these reasons, it w a s necessary t o obtain yaw-control moment on the model from the d i f f e r e n t i a l propeller pitch.
SUMMARY OF RESULTS The r e s u l t s of . t h e f l i g h t t e s t s of the l/g-scale model of a four-propeller tilt-wing transport airplane without a r t i f i c i a l s t a b i l i z a t i o n may be summarized as follows: 1. Hovering-flight tests out of ground e f f e c t showed t h a t basic controls- fixed motions of the model consisted of unstable oscillations i n pitch and r o l l and t h a t the model was neutrally stable i n yaw.
The unstable oscillations were of relatively long period, however, and were very easy f o r the p i l o t t o control.
2. Hovering-flight t e s t s i n ground effect showed t h a t the model had a posi- t i v e ground e f f e c t on l i f t . The pitching o s c i l l a t i o n became l e s s unstable as the model neared the ground and was about neutrally stable when the wheels were j u s t off the ground. The effect of the ground on the r o l l i n g oscillation w a s l e s s pronounced, but the rolling motions became s l i g h t l y e a s i e r t o control as the model neared the ground. The model experienced significant random yaw dis- turbances when hovering near the ground, and there w a s a noticeable reduction i n yaw-control power available, but the yawing motions could be kept under con- t r o l by the p i l o t with suitable attention t o t h e controls.
3 . I n the t r a n s i t i o n range no trouble w a s experienced with e i t h e r longitudi- n a l o r l a t e r a l s t a b i l i t y or control i n l e v e l forward f l i g h t , except t h a t the model had about neutral directional s t a b i l i t y f o r very s m a l l angles of sideslip.
In general, the model had at l e a s t 6 ' descent capability with no adverse e f f e c t s and no noticeable wing s t a l l i n g . Another 3' o r 4' of descent w a s available before stalling caused the flying q u a l i t i e s t o become completely unacceptable.
4. I n a l l f l i g h t regions, the mini" t o t a l control powers found t o be sat- isfactory i n the model f l i g h t t e s t s were less than t h e control power planned f o r the full-scale a i r c r a f t .
Langley Research Center, National Aeronautics and Space Administration, Langley Station, Hampton, V a . , August 5, 1964.
1. P a r l e t t , u s l e P.; and Kirby, Robert H.: Test Techniques Used by NASA f o r Investigating Dynamic S t a b i l i t y Characteristics of V/STOL Models. AIAA Aerodynamic Testing Conf'. , March 1964, pp. 42-49.
2. McKinney, M. 0.; Kirby, R. H.; and Newsom, W. A., Jr.: Aerodynamic Factors t o be Considered i n the Design of Tilt-Wing V/STOL Airplanes. Vertical Take-Off and Landing (VTOL) Aircraft.
Ann. N.Y. Acad. Sci., vol. 107, A r t . 1, I. B. Laskowitz, ed., Mar. 25, 1963, pp. 221-248.
3. Tosti, Louis P. : Flight Investigation of the S t a b i l i t y and Control Charac- t e r i s t i c s of a 1/4-Scale Model of a Tilt-Wing Vertical-Take-Off -and-Landing Aircraft. NASA MEMO 11-4-58L, 1959.
4. Pegg, Robert J.: Summsry of Flight-Test Results of the VZ-2 Tilt-Wing Air- c r a f t . NASA TN D-989, 1962.
5. Newsom, W i l l i a m A., Jr.: Effect of Ground Proximity on the Aerodynamic Char- a c t e r i s t i c s of a Four-Engine Vertical-Take-Off-and-Landing Transport- Airplane Model With T i l t i n g Wing and Propellers. NACA TN 4124, 1957.
6. Schade, Robert 0.: Ground Interference Effects. N A S A TN D-727, 1961.
7. Newsom, W i l l i a m A., Jr.: Flight Investigation of t h e Longitudinal S t a b i l i t y and Control Characteristics of a Four-Propeller Tilt-Wing VTOL Model With a Programed Flap. NASA TN D - 1 3 9 0 , 1962.
8. Schade, Robert 0.; and Kirby, Robert H.: Effect of Wing Stalling i n Transi- t i o n on a l/h-Scale Model of the VZ-2 Aircraft. NASA TN D-2381, 1964.
9. Anon.: Helicopter Flying and Ground Handling Qualities; General Requirements f o r . Military Specification MIL-H-~~OIA, sept. 7, 1961.
T A B L E 1.- GEOKETRIC CHARACTERISTICS OF THE MODEL Fuselage: L e . h . ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.56 Cross-sectional area, maxi", s q f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.01 Height, me;ximum, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.36 Width, maxi", f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.01 wing : Area, s q f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.60 span, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.50 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
A s p e c t r a t i o 8.53 Mean aerodynamic chord, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.90 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Airfoil section NACA 6j3-318 Tip chord, f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.67 Root chord, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.09 Taper r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.61 Sweepback of quarter chord, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.13 Dihedral angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -2.12 Pivot station, percent root chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42.5 Aileron, each: Chord, percent wing chord 25 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Area, s q f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.38 Flap, each: Type . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Double s l o t t e d Chord, percent wing chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47 Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . N l Slat, each: Inboard, 0.45 w i n g semispan t o 0.69 w i n g semispan . . . . . . . . . . . Chord, 0.20 wing chord inboard t o 0.10 wing chord outboard Outboard. 0.85 wing semispan t o 1.00 wing semispan . . . Chord. 0.10 wing chord f u l l length Vertical tail: Basic: Area. s q f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.61 Span, f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.73 Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.87 Airfoil section: Root . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NACA0018 Tip . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . N A C A O O ~ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Tipchord, f t 0.37 Root chord, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.48 Taper r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.25 Sweepback of quarter chord, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26 Rudder : Tip chord, f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.15 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.42 Root chord, f t Span, measured f r o m t i p chord, f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.06 T a i l length, center of gravity t o 0.25 mean aerodynamic chord, f t . . . . . . . . . . . . . . . . . 2.38 Large : Area, s q f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.60 Span, f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.11 Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.71 Tip chord, f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.57 Root chord, f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.89 Taper r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.30 Horizontal tail: Area. s q f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.11 Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.68 A i r f o i l section: Root . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . N A C A O O ~ ~ . . . . . . . . . . . . . . . . . . NACA0012 Tip . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Tip chord. ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.39 Root chord. ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.78 . . . . . . . . . . . . . . . . . . . . 3.46 span. f t . . . . . . . . . . . . . . . . . . . . . . . . . . .
Taper r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.50 Sweepback of quarter chord. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.50 Mean aerodynamic chord. f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.61 f t . . . . . . . . . . . . . . . . . . 2.76 T a i l length. center of gravity t o 0.25 mean aerodynamic chord.
Propellers : kin: Number of blades . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 Diameter. f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.72 T a i l : Number of blades . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Diameter. f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.89 Moment arm. wing pivot t o rotor center. ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.56 . . . - ............ . . . . . . .
TABU 11.- COMPARISON O F AVERAGE MASS CHARACTERISTICS O F MODEL ( S O U P ) AND FULL-SCALE AIRPLANE
Characteristic I Model (scaled up) I Airplane
I I Gross weight. lb . . . .
52. 000 37. 424 Ix. slug-ft2 . . . . . .
307. 000 176 . 430
Iy. slug-ft2 . . . . . . 205. 000 125. ooo
I z . slug-ft2 . . . . . . 418. 000
270. 631
kx. f t . . . . . . . . . 12.2
13.8
ky. f t . . . . . . . . . 11.3 10.4
kz. f t . . . . . . . . . 16.1 15.4
W/S. l b / f t 2 . . . . . .
97.2 70 TABLE =I.- C O M P A F U S O N O F MOEL RATING SYSW WITH COOPER WIT" SYSTEM Flying-model p i l o t - r a t i n g system Cooper p i l o t opinion r a t i n g system Numerical primary
1 Adjective 3perat ing
Description Description mission landed r a t i n g conditions accomplished
I rating l l
Extremely easy t o f l y - requires Excellent, includes optimum Yes C ~ ~ b e
no a t t e n t i o n t o control Normal
Very easy t o f Q - requires p r a c t i - Good, pleasant t o f l y Yes
yes ~ Satisfactory operat ion c a l l y no a t t e n t i o n t o control Easy t o f l y - requires very l i t t l e ' S a t i s f a c t o r y , but with some mildly Ye 6 Yes a t t e n t i o n t o control unpleasant c h a r a c t e r i s t i c s
4 Not d i f f i c u l t t o f l y - requires Acceptable, but with unpleasant Yes Yes
a t t e n t i o n t o control c h a r a c t e r i s t i c s 5 Not too d i f f i c u l t t o f l y - requires Undcceptable f o r normal operation Doubtful Y e 6 Emergency considerable a t t e n t i o n t o control Unsatisfactory operation 6 D i f f i c u l t t o f l ~ - requires almost Acceptable f o r emergency condi- Doubtful Yes constant a t t e n t i o n t o maintain t i o n only1 f l i g h t 7 V e r y d i f f i c u l t t o f l y - requires Unacceptable even f o r emergency N o Doubtful constant a t t e n t i o n t o maintain condition1 f l i g h t
8 Extremely d i f f i c u l t t o fQ - flyable Unacceptable - dangerous No No Unacceptable
only with maximum a t t e n t i o n given No . t o maintain f l i g h t ope rat ion
9 Unflyable - cannot be flown even Unacceptable - uncontrollable No N o
with maximum a t t e n t i o n given t o , maintaining f l i g h t 10 Catastrophic - model destruction Motions possibly violent enough t o No No Cat as t rophic prevent p i l o t escape l F a i l u r e of s t a b i l i t y autgnenter.
I ( a ) Model with iv = Oo and € i f = Oo. L-62-9665 Figure 1.- Photograph of model used in i n v e s t i g a t i o n .
(b) Transition fli@;ht i n Langley m l - s c a l e tunnel. L-63-8475 Figure 1.- Concluded.
fl 7 0 . 6 Diam.
I Ailerons / 1 't, - - - - - I '
- I 8.9 --& I 7.8
I I (a) Three-view sketch of model.
All dimensions a r e in inches.
Figure 2.- Model sketch.
FW-ELEMEX'I COORDINATES Second Element F i r s t Element - Chord Chord Lower Lower Upper Station Ration - -1.20 -1.20 0 0 0 1 . 4 0 -2.89 1 . 4 7 -1.32 1 . 7 1 . 0 2.36 -2.94 2 . 0 -1.49 3.0 1-97 3 . 6 8 -2.28 -1.49 2 . 3 3 5 . 5 3.0 -2.28 8.0 4.38 4 . 0 2 . 5 6 -1.25 -2.06 4 . 5 2 2.71 -0.83 9.5 5.0 12.8 -1.67 2.78 - .56 4.33 6.0 13.0 4.03 -1.46
2.82 - .07
7 . 0 14.2 3.82 -1.31 8.0 2.78 .21 -0.90 2 . 6 5 -40 18.0 3 . 0 7 9.0
- .22
2 . 4 3 23.0 1.75
10.0 - 51
28.0 0.96 .04 2 . 1 5 ~ .56 11.0 0 0 12.0 1.81 67 1.35 .42 13.0 14.0 76 .25 15.4 (b) Typical c r o s s s e c t i o n of wing with double s l o t t e d f l a p showing m a x i " f l a p d e f l e c t i o n and 0 . 2 5 ~ a i l e r o n on second f l a p element.
A l l coordinates i n percent wing chord.
Figure 2.- Concluded.
h -P
I
Model
i
-- Airplane
IO
100 60 40
Wing incidence, Figure 3.- Variation of model and airplane center-of-gravity position with wing incidence.
Flap
- - --- H o r i z o n t a l t a i l
‘i I
I
I 1
I Model
1 and
1 I
u
30 I
1 I
20 1
I I
i i r p l a n e E -
i I
10 1 I
I
I
I
Y 100 80 60 20 0
4 0
Wing incidence, deg Figure 4.- Variation of f l a p d e f l e c t i o n and h o r i z o n t a l - t a i l incidence with wing incidence.
L-64-3008 Figure 5.- Sketch of the setup used for flight tests in the Langley full-scale tunnel.
Actual descent S i m l a t e d descent L i f t L i f t
Dr ag Wind 4 Drat2
Auxiliary j e t t k r u s t 1 - 1
w
Weight Weight Figure 6.- Balance of f o r c e s i n descent and simulated descent conditions.
: . . . i .
a -20 1 5 Figure 7.- Control-fixed pitching oscillation of model in hovering flight out of ground effect.
a -20 rl G
-40
0 1 2 3 6
4 5
T i m e , s e c Figure 8.- Control-fixed r o l l i n g o s c i l l a t i o n of model i n hovering f l i g h t out of ground e f f e c t
-
... ...
r r Wheels touch IIghtly Nose wheel hits hord Nose wheel hits hord -73 Wheels touch IIghtly T C p a l y W L s m d 20- a , - - I ; " Time, sec Figure 9.- Control-fixed pitching motion of model during hovering f l i g h t i n ground e f f e c t .
w 1 . 0 2.0
1.5
2.5 3.0 Figure 10.- Effect of ground proximity on s t a t i c s t a b i l i t y . Wheels touch at h/D = 0.24.
1 . 0 . 8 Constant l i f t Mz Constant p r o p e l l e r
-
r o t a t i o n a l speed M 2,- .6 .4 .2 I L 1 . 0 . 8 * , , , I , , , / / r 7 n , ,,,,, , ,,, ,, , , 7 -h .4 . 2 0 .5 1.0 1.5 2 . 0 2.5 3.0 3.4 Figure 11.- Effect of ground proximity on aileron yaw control effectiveness.
h/D ~ ; i 0.24.
Wheels touch at w N 1.20 1.10 1.00 I 1- -I-- I A
.5 1 . 0
1.5 2.0
2.5
3.0
h b Li Figure 12.- Effect of ground proximity on lift at constant p r o p e l l e r r o t a t i o n a l speed and blade p i t c h angle.
touch at h/D 0.24.
Wheels a ba a t l = l .-- 100 a
B
a a
I
a t F M a a * b) & d -20
-40
0 1 2 3 6 4 5 T l m e , s e c Figure 13.- Control-fixed l o n g i t u d i n a l motions of model i n t r a n s i t i o n f l i g h t range.
w -r
0 Basic model as flown
0 Plus modified rear fuselage lines
0 Plus fuselage-flap juncture taped
a 10
-15 -10 -5 0 5 10 15
Figure 14.- Effect of some model modifications on variation of yawing-moment coefficient with sideslip. i , = OO and a = 00.
-15 .10 .05 cn i, = 100 I I I 1 I I I I I l l -.05 .15 I I I I I I I 1 l I I I I I I I I I I I I I I I I I I I I I I I I L L L L ! I I I I I I I I I I I 0 Baaio model d t h modified rear fuselage lines 0 0 plus fuselage-flap junoture taped .10 * 05
cn
I 1 I I I 1 1 I I 1 1 I 1 I t I i###I -.a5 -15 -10 -5 5 10 Figure 15.- Effect of a l a r g e r v e r t i c a l t a i l on v a r i a t i o n of yawing-moment c o e f f i c i e n t with s i d e s l i p .
w ch
0 Steady-flight ratings
0 Disturbed-flight ratings
@ @ @ @
O r -4 I: Flight-path angle, deg
I
I -
-12 C
c
I I I I I
-16
70 60 50 40 30 20 10
Wing incidence, deg Figure 16.- P i l o t r a t i n g s obtained i n descent t e s t s of a l/g-scale model of a four-propeller t i l t - w i n g t r a n s p o r t .
Unacceptable
0 Unsatisfactory
0 Satisfactory
0 -
Investigation B Range
-4
. . . .
.. .
_ . . .
. . . .
. . , . .
. .
. .
. . ' .
. . .
. .
. . . .
Flight-path -8
angle, deg -12 -16
7 0 60 50 30 20 10
Wing incidence, deg Figure 17.- Descent capability i n t r a n s i t i o n from f l i g h t t e s t s of a l/g-scale model of a four-propeller tilt-wing transport.
w , . . . . . , R o l l
Yaw
-20 -20 6 8 0 2 T i m e , sec i , = 30°.
Figure 18.- Lateral motions of model while attempting smooth flight.
I i a .
Q rl M
B
P
h rl ri ffi -20 M
P
h
a - 20
tl ri rl ffi R o l l -40 M a m rl
z
h a B rl rl [1: -2c y = -100 -4c 0 2 4 T i m e , s e c i, = 30°.
Figure 19.- L a t e r a l motions of model a f t e r a d e l i b e r a t e d i s t u r b a n c e .
Full- scale a i r c r a f t Figure 20.- Longitudinal c o n t r o l power a v a i l a b l e i n excess of t h a t required f o r t r i m on t h e f u l l a i r p l a n e compared with scaled-up model c o n t r o l power required i n t e s t s .
1 . 0 Full- scale a i r c r a f t Scaled-up model a I d
.a
% .6 ief. '
I t i H7-H
rl k
-P 0 4
I rl d p: . 2
2 -
(d a .6 m .11 pedal
. 4
ief. 9 d
I I I I I I I I l l l l l l l l
. 2 k '9 El I k 100 20 0 Figure 21.- L a t e r a l c o n t r o l power a v a i l a b l e on a i r p l a n e compared with scaled-up model c o n t r o l power r e q u i r e d i n t e s t s .
NASA-Langley, 1964 L-4179
A mot ion-picture film supplement L-833 is available
Requests w i l l be f i l l e d i n the order received. on loan.
You will be notified of the approximate date scheduled.
The film (16 mm, 1 7 min, color, s i l e n t ) shows a ver-
NASA TN D-2443 I. Newsom, William A., t i c a l take-off and landing, a slow constant-altitude tran- National Aeronautics and Space Administration. Jr .
s i t i o n from hovering t o forward f l i g h t , hovering flights FLIGHT INVESTIGATION OF STABILITY AND CON- II. Kirby, Robert H.
i n and out of ground effect, control fixed motions i n TROL CHARACTERISTICS OF A l/B-SCALE MODEL I I I . NASA TN D-2443 OF A FOUR-PROPELLER TILT-WJNG V/STOL hovering and forward f l i g h t , and a typical series of TRANSPORT. William A. Newsom, Jr., and f l i g h t s at increasing descent angle.
Robert H. Kirby. September 1964. 41p. OTS price, $1.25. (Film Supplement L-835 available on request. ) Requests f o r the film should be addressed to: (NASA TECHNICAL NOTE D-2443) The tests included hovering flights in and out of Chief, Photographic Division ground effect and level flight and descent tests in NASA Langley Research Center the transition speed range. In each flight condition, Langley Station the stability, controllability, and general flight behavior of the model were investigated. Even Hampton, Va. 23365 though the model was statically and dynamically unstable for many of the flight-test conditions, it could generally be controlled and maneuvered easily.
The descent tests showed that the configu- ration had at least a 6O descent capability with no (over) NASA C U ! P I Date I 'Please send, on loan, copy of film supplement E835 t o I T N D-2443 I I
' Name of organization
I I ~ ~~
; Street number
I ICity and State Zip code I I Attention: Mr.
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