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N A S A TECHNICAL NOTE
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FLIGHT TESTS OF A ONE-MAN
HELICOPTER A N D A COMPARISON OF
ITS HANDLING QUALITIES WITH THOSE
OF LARGER VTOL AIRCRAFT
by Terrell W. Feistel and Fred J. Drinkwater III
Ames Research Center
Moffett Field, Cali$
N A T I O N A L A E R O N A U T I C S A N D SPACE A D M I N I S T R A T I O N W A S H I N G T O N , D. C. OCTOBER 1965 TECH LIBRARY KAFB, NM FLIGHT TESTS O F A ONE-MAN HELICOPTER AND A COMPARISON O F ITS HANDLING QUALITIES WITH THOSE O F LARGER VTOL AIRCRAFT By Terrell W. Feistel and Fred J. Drinkwater I11 A m e s R e s e a r c h C e n t e r Moffett Field, Calif.
NATIONAL AERONAUT ICs AND SPACE ADMlN I STRAT I ON For sale by the Clearinghouse for Federal Scientific and Technical Information
Springfield, Virginia 22151 - Price $1.00
FLIGHT TESTS OF A O N E - i " HELICOPTER AND A COMPARISON OF I T S HANDLING QUALITIES WITH THOSE O F LARGER VTOL AIRCRAFT By T e r r e l l W. F e i s t e l and Fred J. Drinkwater I11 Ames Research Center SUMMARY A l i m i t e d f l i g h t t e s t program has been accomplished with a one-man H i l l e r YROE-1 "Rotorcycle" (gross wt 2 500 l b ) t o help determine c r i t e r i a f o r t h e handling q u a l i t i e s i n hover of WOL a i r c r a f t as a f f e c t e d by gross weight.
The g e n e r a l l y high orders of l o n g i t u d i n a l and lateral c o n t r o l power and damping inherent w e r e found t o be s a t i s f a c t o r y . The high d i r e c t i o n a l c o n t r o l s e n s i t i v i t y , combined w i t h high yaw response i n one d i r e c t i o n , w a s considered p o t e n t i a l l y dangerous. The l a t e r a l c o n t r o l power f o r t h i s c r a f t i s approxi- mately t h e same as t h a t found necessary f o r s a t i s f a c t o r y c o n t r o l with similar damping i n t e s t s of two other VTOL a i r c r a f t w i t h s u b s t a n t i a l l y g r e a t e r gross weight.
INTRODUCTION The N A S A has, i n r e c e n t years, been studying handling q u a l i t i e s c r i t e r i a f o r V/STOL a i r c r a f t (ref. 1 ) . A major question i s , how do s a t i s f a c t o r y and u n s a t i s f a c t o r y l i m i t s f o r hovering c o n t r o l power and damping vary with s i z e and gross weight? One form of s c a l i n g c r i t e r i a i s presented i n reference 2.
Only l i m i t e d f l i g h t v e r i f i c a t i o n of these c r i t e r i a i s a v a i l a b l e , c h i e f l y w i t h vehicles i n t h e 3000-4000 pound g r o s s weight category (see, e.g., ref. 3 ) .
The H i l l e r YROE-1 Rotorcycle, with a gross weight of approximately 500 pounds, i s a t the bottom end of the weight spectrum f o r manned a i r c r a f t and an order It w a s s e l e c t e d f o r of magnitude away f r o m t h e X-14A (used i n r e f . 3 ) .
i n v e s t i g a t i n g c o n t r o l power requirements a t low gross weights i n hope t h a t , thereby, some l i g h t would be shed on the influence of s i z e and weight.
DESCRIPTION O F TEST A R T I C U The H i l l e r YROE-1 Rotorcycle ( f i g . 1) w a s o r i g i n a l l y designed f o r t h e Armed Services as a simple, c o l l a p s i b l e , one-man h e l i c o p t e r f o r observation Figure 2'shows a three-view sketch of t h e vehicle. A s and l i a i s o n purposes.
flown, i t s gross weight w a s 515 pounds. The power p l a n t i s a 4-cylinder, The craft i s described i n d e t a i l i n refer- 2-cycle, Nelson engine of 43 hp.
ences 4 and 5 and results of previous Navy f l i g h t tests are given i n refer- ences 6 and 7.
- _ - F ~~ .___- - - 'Figure 2 was supplied by t h e H i l l e r A i r c r a f t COT., Inc.
I
l l 1 1 l 1 l 1 1 l l I l 1 l 1 1 1 1 1 1 1 l I II I1 I I
For the NASA f l i g h t t e s t s a small instrumentation package w a s hung i n a box underneath t h e p i l o t as shown i n f i g u r e 1. The package contained a high- frequency t r a n s m i t t e r t h a t telemetered information on three channels, and a s i n g l e - a x i s rate-measuring gyro which could b e o r i e n t e d along any one of the three axes. A potentiometer w a s a l s o included f o r measuring t h e p o s i t i o n of t h e c o n t r o l being considered and a b u t t o n on t h e c o n t r o l s t i c k allowed t h e p i l o t t o s i g n a l the start of a maneuver.
M E T H O D O F DATA mDUCTION The maneuver f o r obtaining t h e c o n t r o l power-damping d a t a consisted of a c o n t r o l reversal input t o t h e c o n t r o l a f f e c t i n g t h e axis being considered ending i n a f i x e d c o n t r o l d e f l e c t i o n held f o r 1 - 2 seconds. Thus, t h e f i r s t d e r i v a t i v e of the r e s u l t a n t angular v e l o c i t y abaut t h e given axis, as t h i s v e l o c i t y passes through zero a f t e r the reversal, represents the angular a c c e l - e r a t i o n corresponding t o t h e c o n t r o l d e f l e c t i o n . The m a x i m excursion of t h e angular v e l o c i t y ( w i t h t h e c o n t r o l s t i l l h e l d f i x e d ) , when compared t o t h e previously determined angular acceleration, i n d i c a t e s t h e approximate v e l o c i t y damping, 1/~, about t h e axis. See reference 8 f o r an a n a l y s i s of t h i s method.
The angular a c c e l e r a t i o n corresponding t o t o t a l c o n t r o l d e f l e c t i o n w a s d e t e r - mined by p l o t t i n g a c c e l e r a t i o n s measured a g a i n s t percent d e f l e c t i o n and fair- i n g a s t r a i g h t l i n e through a l l t h e p o i n t s (assuming a l i n e a r v a r i a t i o n of i n i t i a l angular a c c e l e r a t i o n with c o n t r o l d e f l e c t i o n ) .
The p i l o t r a t i n g s were based on t a s k s described i n t h e P i l o t Comments s e c t i o n .
RESULTS AND DISCUSSION Since t o t a l c o n t r o l power required f o r a given t a s k includes t h a t f o r c o r r e c t i n g d i s t u r b i n g inputs, it w i l l depend on t h e type of VTOL a i r c r a f t because of inherent d i f f e r e n c e s i n gust s e n s i t i v i t y , e t c . I n s p i t e of t h i s it i s of i n t e r e s t t o examine c o n t r o l power requirements f o r a wide range of VTOL aircraft t o observe any gross t r e n d of t h e e f f e c t of s i z e .
Table I i s a summary of t h e p e r t i n e n t parameters determined. It shows m a x i m u m c o n t r o l power ( i n terms of i n i t i a l angular a c c e l e r a t i o n ) , angular rate damping ( i n terms of t h e r e c i p r o c a l of the t i m e constant, l / - i - ) , c o n t r o l s e n s i - t i v i t y ( i n terms of i n i t i a l a c c e l e r a t i o n p e r inch of c o n t r o l d e f l e c t i o n ) , and t h e p i l o t r a t i n g f o r the visual hovering t a s k ( f o r both m a x i m u m c o n t r o l power and s e n s i t i v i t y , where a v a i l a b l e ) on t h e Cooper Scale ( t a b l e I1 and ref. 9 ) f o r each of t h e three axes (two p i l o t r a t i n g s are shown; p i l o t A being t h e p r o j e c t p i l o t and p i l o t B a v i s i t i n g N A S A p i l o t who made only one f l i g h t ) .
For the d i r e c t i o n a l case values are shown f o r r i g h t yaw only. A l s o shown are t h e implied values of t h e c o n t r o l power, damping, and s e n s i t i v i t y c a l l e d out i n t h e proposed V/STOL s p e c i f i c a t i o n s ( r e f . 2) . These have been converted from t h e response values l i s t e d by assuming a "step" input t o t h e c o n t r o l .
For comparison similar d a t a are shown, i n parallel grouping, f o r t h e minimal s a t i s f a c t o r y (P.R. = 3.5) r a t i n g i n t h e v a r i a b l e s t a b i l i t y X-14A (used i n ref. 3 ) . A l s o shown i n t h e "damping" column are t h e nominal moments of i n e r t i a of t h e two c r a f t i n slug-ft2.
Lateral Characteristic s Figure 3 shows a p l o t of t h e i n i t i a l a c c e l e r a t i o n i n roll f o r f u l l con- t r o l d e f l e c t i o n (i.e., c o n t r o l power) and f o r one inch of c o n t r o l t r a v e l (i.e., s e n s i t i v i t y ) versus gross weight f o r four vehicles: t h e YROE-1 helicopter ( W = 515 l b ) , t h e X-14A deflected j e t VTOL ( W = 3880 l b , r e f . 3 ) , t h e Hawker P-1127 deflected j e t VTOL ( W = 12,500 l b ) , and t h e XC-142A t i l t - w i n g VTOL t r a n s p o r t ( W = 37,500 l b , r e f . 1 0 ) . Data f o r t h e latter two were supplied by t h e manufacturer. P i l o t r a t i n g s f o r t h e v i s u a l hovering t a s k , where a v a i l a b l e , a r e shown i n parentheses next t o t h e data points; f o r t h e YROE-1, t h e r a t i n g s of t h e p r o j e c t p i l o t only a r e shown. The l a c k of v a r i a t i o n with gross weight of c o n t r o l power required t o obtain a s a t i s f a c t o r y (P.R. = 3-1/2) r a t i n g f o r t h i s important "X" a x i s i s of i n t e r e s t . For t h e YROE-1, t h e p i l o t r a t i n g of u n s a t i s f a c t o r y f o r s e n s i t i v i t y i n roll (and a l s o i n p i t c h ) w a s given because of t o o l i t t l e s e n s i t i v i t y ( t o o much s t i c k t r a v e l , +7 i n . i n roll). P i l o t r a t i n g s f o r s e n s i t i v i t y i n t h e P-1127 and f o r t h e XC-142A a r e not available.
Reference 11 w a s used t o derive a p i l o t r a t i n g f o r t h e s e n s i t i v i t y of t h e X-14A i n roll; t h e s e n s i t i v i t y shown corresponds t o t h e 3-1/2 boundary f o r c o n t r o l power.
Figure 4 i s another p l o t showing handling q u a l i t i e s information i n roll, The " s a t i s f a c t o r y " (P.R. = 3-l/2) and "acceptable" with some of t h e same data.
(P.R. = 6-1/2) boundaries f o r lateral c h a r a c t e r i s t i c s axe shown as determined The boundaries are p l o t t e d with with t h e variable s t a b i l i t y X-14A (ref. 3 ) .
t o t a l c o n t r o l power as t h e abscissa and r a t e damping ( t h e r e c i p r o c a l of t h e Superimposed i s a point showing t h e charac- t i m e constant) as t h e ordinate.
t e r i s t i c s of t h e YROE-1 as determined by t h e subject tests; it i s seen t o pos- sess, with a p i l o t r a t i n g of 3, approximately t h e same control power and damp- ing i n roll as w a s required by t h e X-14A f o r a s a t i s f a c t o r y p i l o t r a t i n g . The values shown f o r t h e P-1127 correspond t o t h e configuration flown by a N A S A p i l o t when t h e r a t i n g of 3-l/2 w a s assigned. The values f o r t h e X C - 1 4 u l are f o r t h e unaugmented configuration and a r e estimates only.
Longitudinal C h a r a c t e r i s t i c s It can be seen Figures 5 and 6 show t h e handling q u a l i t i e s i n p i t c h .
t h a t t h e YROE-1 ( w i t h a P.R. of 2-3) possesses much higher values of c o n t r o l power and damping about t h e Y a x i s than were necessary f o r s a t i s f a c t o r y (P.R. = 3-l/2) c h a r a c t e r i s t i c s i n t h e X-14A or t h e P-1127. This combination of control power and damping w a s t o o high t o be evaluated i n t h e X-14A, but it i s s i g n i f i c a n t that t h e p i l o t r a t i n g i n d i c a t e s l i t t l e improvement over t h e r a t i n g s obtained a t t h e lower l e v e l s of c o n t r o l power and damping along t h e 3.5 boundary of reference 3. A s i n t h e l a t e r a l case, t h e longitudinal c o n t r o l s e n s i t i v i t y w a s r a t e d at 5 because of t h e l a r g e (+_8 i n . ) s t i c k t r a v e l . The values shown f o r t h e XC-142A a r e estimates f o r t h e unaugmented configuration.
Directional C h a r a c t e r i s t i c s Because of t h e unusual circumstances involved, no comparison p l o t s a r e shown f o r t h e d i r e c t i o n a l c h a r a c t e r i s t i c s . Too much c o n t r o l power _ . . . . .
I l 1 l 1 1 l 1 l 1 l 1 l 1 1 1 l 1 1 1 1 1 I I I I l l l l l I
(approximately 6 radians/sec2 i n hover near sea l e v e l ) i s a v a i l a b l e t o t h e r i g h t (aiding r o t o r torque) along with a n extremely high s e n s i t i v i t y (approxi- mately 3 radians/sec2/in. corresponding t o 2 i n . pedal t r a v e l ) , which were given p i l o t r a t i n g s of 6 and 7, respectively. The p i l o t must be highly compe- t e n t t o f l y t h e vehicle successfully because of t h i s high s e n s i t i v i t y and c o n t r o l power i n yaw. I n t h e opposite d i r e c t i o n ( t o t h e l e f t , countering r o t o r torque) no accurate measurements could be taken since c o n t r o l power i s marginal and v a r i e s considerably with f l i g h t condition as a r e s u l t of t h e varying power input t o t h e r o t o r accompanied by t h e varying t a i l r o t o r thrust required and a v a i l a b l e ( i n ref. 8, f i g . 2, it i s shown t h a t , f o r density a l t i - tudes i n excess of approximately 3000 f t , i n s u f f i c i e n t d i r e c t i o n a l c o n t r o l e x i s t s t o counteract r o t o r torque).
PILOT COMMENTS The p i l o t r a t i n g s of control power, s e n s i t i v i t y , and damping provided i n t h i s r e p o r t a r e based on t h e vehicle c h a r a c t e r i s t i c s when hovering and maneuvering at l o w speeds i n a r e l a t i v e l y confined area.
Lateral-control power i s not t h e same f o r l e f t and r i g h t inputs, and r o l l - p i t c h cross coupling e x i s t s f o r abrupt c o n t r o l displacements. P i t c h and r o l l c y c l i c c o n t r o l displacements a r e excessive and t h e low c o n t r o l s e n s i t i v - i t y contributes t o a f e e l i n g of sluggish p i t c h and roll response, It a l s o f e e l s as though t h e r e i s a delay i n t h e c o n t r o l response from t h e time a step input i s applied t o t h e time t h e response i s f e l t . F u l l lateral c o n t r o l w a s often used i n roll r e v e r s a l maneuvers about t h e hover condition; however, p r e c i s i o n hovering over a spot w a s accomplished wit'n very s m a l l l a t e r a l - c o n t r o l inputs.
Longitudinal c o n t r o l power w a s never l i m i t i n g i n any maneuver. F u l l c o n t r o l w a s used for t h e most abrupt quick stops, b u t , as w a s noted f o r t h e w a s a l a g i n t h e response of t h e helicopter t o abrupt l a t e r a l control, t h e r e c o n t r o l inputs. These effects, which a r e s i m i l a r t o those of t h e l a r g e r H i l l e r 12E, a r e reportedly due t o t h e c h a r a c t e r i s t i c s of t h e servo-paddle- r o t o r cyclic-control system. The p i t c h cyclic-control displacement i s uncom- f o r t a b l y l a r g e , p a r t i c u l a r l y f o r an overhead c y c l i c s t i c k . There w a s noobjec- tionable f r i c t i o n i n t h e c y c l i c c o n t r o l and t h e f o r c e s were very d e s i r a b l e .
Adequate c o n t r o l centering was available, and t h e r o t o r feedback through t h e c y c l i c s t i c k w a s only noticed when abrupt c o n t r o l inputs were used.
P i t c h and roll damping appeared high and considerable s t a b i l i t y i n t e r m s of a roll o r p i t c h r e s t o r i n g moment as a function of forward or sideward speed w a s present.
Yaw c o n t r o l power during hovering w a s high t o t h e r i g h t but just adequate t o t h e l e f t a t normal rpm. It w a s easy t o l o s e a l l d i r e c t i o n a l c o n t r o l power t o t h e l e f t i f t h e r o t o r rpm w a s allowed t o decay t o t h e lower r o t o r speed normal operating l i m i t . Yaw c o n t r o l w a s t o o s e n s i t i v e i n normal hover and w a s considered t o be dangerous for general use because of t h e rocker-plate type of rudder pedals and t h e very high pedal s e n s i t i v i t y , Yaw rate damping appeared t o be high enough and usable yaw r a t e s were not l i m i t e d by t h e rate damping or c o n t r o l power a t high r o t o r rpm.
I n general, t h e r e w a s a tendency t o operate t h i s s m a l l h e l i c o p t e r i n a much t i g h t e r p a t t e r n than even t h e UH-12E ( t h r e e p l a c e , 2800 l b gross w t ) Transitions t o and from a hover were easily done a t high rates h e l i c o p t e r .
and t h e s m a l l s i z e of t h e h e l i c o p t e r minimized t h e judgment needed t o keep a safe distance from obstacles. Operating and observing t h i s small h e l i c o p t e r i n d i c a t e s t h a t it w a s being flown d i f f e r e n t l y than a f l y i n confined areas h e l i c o p t e r of even 2 8 0 0 - p o ~ dgross weight. Turns and t r a n s i t i o n s t o and from hover were done much quicker than i s normally done with t h e l a r g e r helicopters.
The cyclic-control power and rate damping d i d not l i m i t t h e maneuverability of t h e YROE-1 i n and about t h e v i s u a l hover condition. The high yaw c o n t r o l sen- s i t i v i t y required more than normal p i l o t a t t e n t i o n and considerable familiar- i z a t i o n time.
CONCLUDING REMEiRKs The m o s t s i g n i f i c a n t d a t a obtained i s t h a t representing t h e l a t e r a l c h a r a c t e r i s t i c s . This i n d i c a t e s t h a t approximately t h e same l a t e r a l c o n t r o l i s required f o r t h i s vehicle as f o r those of much higher gross weights power t o achieve a s a t i s f a c t o r y p i l o t r a t i n g . The i n d i c a t i o n would seem t o be t h a t minimum c o n t r o l power requirements should be based p r i m a r i l y on t h e t a s k t o be performed r a t h e r than on t h e gross weight o r s i z e , as such.
The d a t a obtained about t h e other two axes i s l e s s conclusive. Appar- e n t l y , t h e high c o n t r o l power a v a i l a b l e longitudinally i s i n e f f e c t i v e because of t h e l a r g e s t i c k movements necessary with consequent low s e n s i t i v i t y .
Directionally, t h e low c o n t r o l power i n t h e d i r e c t i o n opposing r o t o r torque and high c o n t r o l power i n t h e opposite d i r e c t i o n , combined w i t h extremely high c o n t r o l s e n s i t i v i t y , a r e e s s e n t i a l l y p e c u l i a r t o t h i s vehicle and make t h e r e s u l t s inapplicable i n any general sense.
It i s t o be noted t h a t undue emphasis should not be placed on making comparisons of t o t a l c o n t r o l power requirements between d i s s i m i l a r types of VTOL vehicles ( i . e . , h e l i c o p t e r , deflected j e t , tilt wing, e t c . ) , because of inherent differences i n self-disturbing c h a r a c t e r i s t i c s , ground e f f e c t s , gust s e n s i t i v i t y , t r i m requirements, e t c . The comparisons made here a r e presented p r i m a r i l y t o provide a convenient cataloging of a v a i l a b l e d a t a on VTOL air- c r a f t covering a wide range of gross weights and t o p o i n t out t h a t no gross t r e n d of varying c o n t r o l power requirements with increasing weight i s obvious.
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Rolls, Stewart L.; Drinkwater, Fred J., 111; and I n n i s , Robert C . : 11.
E f f e c t s of L a t e r a l Control C h a r a c t e r i s t i c s on Hovering a J e t L i f t VTOL A i r c r a f t . N A S A TN D-2701, 1965.
SUMMARY O F THE PERTINENT PARAIVlETERS DETERMINED TAB= I.- I P i l o t r a t i n g s Control power, Damping, Sensitivity, -radians/sec2 -1/T = l / s e c -radians/secZ/in. , Control 'ens'- i power t i v i t y Mode A/C I I' Flight AGARD Flight t e s t spec. A B A B value I ( r e f . 2) value ' ( r e f . 2) value ,
1 . 7 1 3.9 0.24 ' 1 . 3 3 4 5 4 '
YROE- 1 T - L ^ _ ^ ^ 1 I (-In'? 1170) i satisfactory ) I
. 6 i2-1/2 5 ' E 2.0
2.4 , 'e .25
( t o o low] YROE-1 I ( - I ~ z 80) l Longitudinal/ (-Pitch) ' X- 1 4 A .11 I -25 (3-1/2) I I
--
I Directional
t o r que high) ( -Yaw)
X-14A - 5 .7 1.0 2.4 .17 .23 (3-1/2) (n.a.>
(minimal s a t i s f a c t o r y ) ("Izz 3 2920)
I
TABU 11.- PIL9T OPINION RATING SCHEDULE ~~~ Primary Can b e mission Description r a t i n g r a t i n g landed ac c omp 1 i shed Excellent, includes optimum Yes Yes I 1 2 Good, pleasant t o f l y Yes Yes Unsatisfactory Satisfactory, but with some mildly operat i o n
unpleasant c h a r a c t e r i s t i c s Yes I Yes
Acceptable, b u t with unpleasant
l 4
c h a r a c t e r i s t i c s Yes Yes Unacceptable f o r normal operation Doubtful Yes Acceptable f o r emergency condition 1 I only* Doubtful Yes
I
Unacceptable even f o r emergency
I I. 7
condition* No 1 Unsatisfactory '
Unacceptable - dangerous
operation Unacceptable- unc o n t r o l l a b l e ?ailwe of a s t a b i l i t y augmenter A-31028 Figure 1.- YROE-1 Rotorcycle in hovering flight.
\D Figure 2.- Three-view drawing of test vehicle.
@ Full deflection I (- Total control power) One inch ( w Control sensitivity) I
i I
I i-
I R O E - I I I 1 ~ x - 1 4 ~ I
i I
P.R.- 4
t
RR .N 5
i (Too low)
i
I o2 2 4 6 8 IO3 2 4 6 8 lo4 2 4 6 8 1
Gross weight, W, Ib Figure 3 .- Lateral control characteristics ( - visual hovering task).
P P -4 - 3 - 2 - I I 0 I 2 3 4 5
L a t e r a l c o n t r o l p o w e r , 6, r a d i a n s / s e c
Figure 4.- Lateral handling characteristics.
I
I I 0 Full deflection (- Total control power) (u E One inch a , U J (-Control sensitivity)
>
C I 0 2
E
L :8 ..
C .- .I-
E
a , F Y R O E - I ' I Z3 P.R.- 5 (Too low) I 1 1 1 1 IO2 2 4 6 8 IO3 2 4 6 8 IO' 2 6 8 1 Gross weight, W , Ib Figure 5.- Longitudinal control c h a r a c t e r i s t i c s ( - v i s u a l hovering t a s k ) .
- 2.4
I I I f R O E - I ( P . R . m 2 - 3 ) -
- 2.0
- 1.6
- 1.2
I m P - 1 1 2 7 ( P . R . - 3 2 ) - U7 ( E s t . d a m p i n g ) I - X - 1 4 A ( R e f . 3 ) I I U n a u g .
@ X C - 1 4 2 A ( E s t . )
- -
(P.R.I.. 4)
2 .o
0 .4 .8 I .2 I . 6 ..
L o n g i t u d i n a l control p o w e r , 8 , r a d i a n s / s e c
Figure 6 . - Longitudinal handling characteristics.
NASA-Langley, 1965 A-857 “The aeronautical and space activities of the United States shall be
conducted so as to contribute . . . to the expansion of human RnowI-
edge of phenomena in the atmosphere and space. The Administration shall provide for the widest practicable and appropriate dissemination of information concerning its activities and the results thereof .” -NATIONAL AERONAUTICS A N D SPACE ACT OF 1958
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