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N A S A TECHNICAL N O T E
N A S A T N D-2700
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A SYSTEM FOR VARYING
THE STABILITY A N D CONTROL
OF A DEFLECTED-JET
FIXED-WING VTOL AIRCRAFT
by Frank A. Pauli, Daniel M . Hegarty,
and Thomas M . Valsh
Ames Resemch Center
Mofett 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. M A R C H 1965 TECH LIBRARY KAFB, NM ' O
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0077747 A SYSTEM FOR VARYING T H E STABILITY AND CONTROL OF A D E F L E C T E D - J E T FIXED-WING V T O L AIRCRAFT By F r a n k A. P a u l i , D a n i e l M. H e g a r t y , a n d T h o m a s M. W a l s h A m e s R e s e a r c h C e n t e r Moffett Field, Calif.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For s a l e by the O f f i c e of T e c h n i c a l Services, Department of Commerce, Washington, D.C. 20230 -- P r i c e $2.00 A SYSTEM FOR VARYING THE STABILITY APJD CONTROL OF A DEFLECTED-JET FIXED-WING VTOL AIRCRAFT By Frank A. Pauli, Daniel M. Hegarty, and Thomas M. Walsh Ames Research Center Moffett Field, Calif.
SUMMARY A system to provide variable control power and augmented stability for a On the Bell X-14A aircraft the exhaust hovering VTOL aircraft is described.
of the two jet engines is deflected to provide lift for hovering flight.
Bleed air is taken from the engines and used for reaction control at the wing tips and tail for attitude control while hovering. Two sets of reaction The original set is mechanically nozzles are used on the modified X-14A.
actuated by the pilot; the other set, which uses electric servo-driven noz- les, was added to generate variable-stability control moments about all axes, Variable-stability modes provided are rate damping, cross-coupling cancella- tion, augmented pilot control, and stiffness with maneuverability cutout.
The control system for one axis is outlined, and criteria for dividing bleed air between the two sets of nozzles are presented. The development of the nozzles for constant flow and pilot safety controls are discussed.
INTRODUCTION In aircraft handling qualities research,variable-stability aircraft make possible the exploration of a wide range of the fundamental aircraft param- eters, such as damping and control power about each axis. Such a research aircraft enables the pilot to evaluate combinations of variables under real- istic flight conditions involving specific tasks. Investigations of new handling qualities concepts, exploration of characteristics of aircraft in the design stage, or improvements of existing aircraft may all be accomplished.
The technique for varying control power and dynamic response depends on supplying commands to the control surfaces, or auxiliary control devices, in addition to the pilot's normal commands. These additional commands are deter- mined on the basis of both pilot command and aircraft motions. Servo- mechanisms are convenient for introducing these commands into the control system. Variable -stability systems are quite flexible since commands can easily be varied in amplitude, mixed in various 'proportions, reversed, or even adjusted to put the aircraft into an unstable condition.
-stability aircraft is particularly valuable in The concept of variable the solution of VTOL stability and control problems. The inherent aerodynamic I damping of t h e VTOL a i r c r a f t tends toward zero when f l i g h t v e l o c i t y approaches zero and hovering i s reached. Since c o n t r o l system research has shown t h a t t h e a b i l i t y of a p i l o t t o c o n t r o l a system depends on t h e damping and c o n t r o l parameters, t h e choice of these parameters f o r VTOL a i r c r a f t i s exceedingly important, first, t o a s s u r e good handling q u a l i t i e s , and second, t o l i m i t t o t a l power requirements. Reference 1 r e p o r t s t y p i c a l r e s u l t s obtained with an a i r c r a f t equipped with v a r i a b l e - s t a b i l i t y systems t h a t allowed t h e c o n t r o l power and damping t o be varied about a l l t h r e e axes. The p r e s e n t r e p o r t describes t h e o r i g i n a l a i r c r a f t c o n t r o l system, t h e v a r i a b l e - s t a b i l i t y system, t h e servo system design with p a r t i c u l a r emphasis on t h e motorized nozzle and s a f e t y subsystem, simulation s t u d i e s , and system performance.
DESCRIPTION O F THE ORIGINAL AIRCRAFT CONTROL SYSTEM The B e l l X - 1 4 VTOL ( f i g . l ) , a fixed-wing, jet-propelled, d e f l e c t e d - j e t a i r p l a n e , w a s t h e t e s t bed f o r t h e research program ( s e e r e f . 2 f o r a complete a i r c r a f t d e s c r i p t i o n ) . The exhaust from t h e j e t engines passes through cas- cade d i v e r t e r s which enable t h e p i l o t t o s e l e c t any condition between horizon- t a l and v e r t i c a l t h r u s t , or t o make a t r a n s i t i o n from one t o t h e o t h e r i n t h e air. During hovering f l i g h t , t h e p i l o t c o n t r o l s t h e a t t i t u d e of t h e a i r c r a f t by one of two s e t s of a i r j e t nozzles a t t h e t a i l and wing t i p s ; t h e other s e t of nozzles w a s added t o provide t h e v a r i a b l e - s t a b i l i t y c h a r a c t e r i s t i c s . The a i r f o r a l l nozzles i s bled from t h e compressors of t h e t u r b o - j e t engines, thus d i v e r t i n g t h r u s t from t h e propulsion system.
The nozzles used by t h e p i l o t a r e mechanically connected t o h i s c o n t r o l s ( t h e s t i c k and t h e p e d a l s ) . The p i t c h nozzle i s on t h e t a i l and has e x i t a r e a s top and bottom. Changing t h i s d i f f e r e n t i a l a r e a produces a p i t c h i n g moment; t h e t o t a l nozzle e x i t a r e a i s a constant. The nozzles on t h e wing Changing t h e difference between the t i p s a r e used f o r r o l l and yaw c o n t r o l .
l e f t nozzle e x i t a r e a and t h e r i g h t nozzle e x i t a r e a generates a r o l l i n g moment. A yawing moment i s created by r o t a t i n g t h e t h r u s t vectors from t h e l e f t and r i g h t nozzles i n opposite d i r e c t i o n s about t h e l a t e r a l a x i s of the a i r c r a f t . The t o t a l nozzle e x i t a r e a of t h e roll-yaw system i s constant.
The b a s i c X-14 a i r c r a f t has gyroscopic cross coupling, as t h e r e s u l t of t h e angular momentum of t h e engine r o t o r s , between t h e p i t c h and yaw axes of t h e a i r c r a f t . For example, a pitch-up motion causes a l e f t yawing moment proportional t o t h e p i t c h i n g r a t e of t h e a i r c r a f t . From a handling q u a l i t i e s standpoint, t h i s coupling limits t h e p i t c h r a t e a t which yawing can be con- t r o l l e d .
VARIABLE -STABILITY SYS'IIEM The first research program using t h e X-14 a i r c r a f t w a s d i r e c t e d a t examining t h e c o n t r o l power and damping r e l a t i o n s h i p i n t h e hovering mode t o define boundaries of good, acceptable, and unacceptable c o n t r o l ( r e f . 3).
To c a r r y out t h a t research program, t h e following system requirements w e r e established: Change t h e e x i s t i n g maximum obtainable a c c e l e r a t i o n proportions from a.
10:3.2:2.8 f o r roll, p i t c h , and yaw, r e s p e c t i v e l y , t o 10:5:2.8.
Provide maximum c o n t r o l and damping c a p a b i l i t y c o n s i s t e n t with t h e b .
c o n s t r a i n t s of weight and safety.
Keep added weight low t o conserve t h e a l r e a d y s h o r t f l i g h t time (time c.
of hovering f l i g h t i s now approximately 15 minutes).
d.
Keep added system independent of t h e p i l o t ' s b a s i c hovering c o n t r o l system. This meant t h a t t h e v a r i a b l e - s t a b i l i t y system operated i n p a r a l l e l with t h e p i l o t ' s b a s i c hovering c o n t r o l system.
e. Provide t h e a b i l i t y t o vary t h e c o n t r o l and damping parameters i n t h e various modes of operation.
To enable t h e s e requirements t o be met, t h e two o r i g i n a l engines w e r e replaced with General E l e c t r i c J - 8 5 - 5 A engines t o provide a g r e a t e r amount of bleed a i r f o r r e a c t i o n c o n t r o l and t o give 25 percent more t h r u s t with about 400 pounds less weight. Using t h e s e engines a l s o r e s u l t e d i n less angular momentum and associated cross coupling.
The system added t o t h e b a s i c a i r c r a f t w a s designed so t h a t c o n t r o l power and damping could be varied independently f o r a l l t h r e e axes and t o provide c o n t r o l cross coupling i f desired. The p i l o t w a s given c o n t r o l of a l l modes and had t h e option of r e v e r t i n g t o t h e b a s i c a i r c r a f t a t any time. For s a f e t y , e l e c t r i c a l l y operated air shutoff valves were p u t i n t h e added system.
The augmented system provided c o n t r o l moments p r o p o r t i o n a l t o t h e p o s i - t i o n of t h e p i l o t ' s c o n t r o l s t i c k and pedals, b u t t h e moments were indepen- The damping moments were l i n e a r d e n t l y a d j u s t a b l e i n magnitude and d i r e c t i o n .
functions of t h e a i r c r a f t angular rates.
A s t i f f n e s s generator s t a b i l i z e d t h e a i r c r a f t about a given a t t i t u d e during r a p i d disturbances, thus reducing t h e p i l o t ' s e f f o r t t o maintain a given a t t i t u d e i n t h e presence of gusts. A maneuverability cutout automati- c a l l y d i s a b l e s t h e s t i f f n e s s c i r c u i t whenever t h e p i l o t moves t h e c o n t r o l s more than a p r e s e t percentage of t h e maximum c o n t r o l motion. This allows t h e a i r c r a f t t o b e maneuvered e a s i l y during t h e hovering task.
The s a f e t y of t h e a i r c r a f t and p i l o t w e r e major considerations i n t h e design. Because of i t s a d j u s t a b l e c h a r a c t e r i s t i c s , a v a r i a b l e - s t a b i l i t y sys - t e m i s b a s i c a l l y more complex than a normal a i r c r a f t c o n t r o l system and t h i s i n h e r e n t l y leads t o g r e a t e r chance of f a i l u r e .
S p e c i a l s a f e t y considerations w e r e required because t h e p i l o t may examine nearly uncontrollable conditions t o e s t a b l i s h boundaries of c o n t r o l l a b i l i t y .
I n the event t h e s e boundaries are exceeded, it i s necessary t h a t t h e p i l o t b e able to return the aircraft to a safe flight condition quickly and without disturbance. It is also important that the transition from the normal air- craft to the variable-stability operation be smooth and that inputs to the variable-stability system not be sensed through force or motion at the pilot's controls.
Since the bleed air of the jet engines,is used as the source of power for the pilot's basic control system, it is advantageous to use it as a source of power for the variable-stability system also. The characteristics of the jet engines made it essential that the total amount of bleed air be constant.
The air flow through the basic control nozzles is constant; therefore, the added flow for the variable-stability nozzles also has to be constant. This assures constant flow conditions in the air ducts.
Check valves in the bleed air discharge lines from each engine prevented the bleed air from reversing the rotation of an engine if it should fail.
The size of these check valves was increased to accommodate an increase in air flow.
Separate air ducts were provided for the variable-stability nozzles so that air to these nozzles could be shut off. This assures that a failure in the variable-stability system cannot offset any of the pilot's basic control moments. It was decided that the basic control nozzles should produce at least 10 percent more moment in each axis than the variable-stability control system. An automatic failure detection circuit also was included in the sys- tem. A button was provided on the stick grip so that the pilot can immedi- ately transfer from the variable-stability mode to the normal aircraft configuration.
SYSTEM DESIGN Bleed air was selected as the source of reaction power forthe X-14A servo system for several reasons. (1) No weight need be added for the source although additional ducting was required. (2) Any other equivalent source, such as air bottles or reaction devices, would necessitate too large a weight addition. (3) A constant ratio of maximum acceleration for the pilot control to the maximum acceleration for the variable-stability system was assured for any axis, thus giving the pilot the same percentage of override capability no matter what engine speed was used or what failure might occur in the air supply The total air-flow rate for both engines was 85.0 pounds per second. A constant 10 percent, or 8.5 pounds per second, of the total engine air was to be bled off for all reaction control. This was greater than recommended by the manufacturer, but it was estimated to be permissible for short flight times. Engine data for 8 . 5 pounds per second of bleed air gave a pressure of 102.9 psia and a temperature of 5 1 6 ' F. Pressure drops through the engine bleed ports, check valves, shut-off valves, motorized valves and ducting, plus the drop necessary for establishing a flow of about 300 feet per second, ... .
The t o t a l r e s u l t e d i n an a v a i l a b l e nozzle pressure of 76.0 p s i a a t 497 F.
e x i t a r e a of t h e nozzles w a s computed t o be 7.17 square inches (appendix).
The t o t a l r e a c t i o n f o r c e i s 500 pounds f o r t h i s air pressure and e x i t a r e a of t h e nozzles (appendix). This f o r c e w a s d i s t r i b u t e d among a l l t h e r e a c t i o n To change t h e maximum nozzles t o meet t h e previously s t a t e d requirements.
force of any nozzle, i t s e x i t a r e a must be changed, with an o f f s e t t i n g change i n t h e a r e a of other nozzles.
To maximize torques from t h e a i r f o r c e s a v a i l a b l e , t h e c o n t r o l nozzles were located on long moment arms. Because of space and ducting problems, t h e v a r i a b l e - s t a b i l i t y roll nozzles were placed on t h e wing t i p s , while t h e p i t c h and yaw nozzles were placed a t t h e t a i l . Only one independent f o r c e e x i s t s f o r each of t h e p i l o t ' s roll-yaw nozzles. All t h e v a r i a b l e - s t a b i l i t y nozzles are separate s o t h a t no inadvertent cross-coupling e f f e c t s w i l l be introduced.
The l o c a t i o n and magnitude of individual nozzle f o r c e s a r e shown i n f i g u r e 2.
The corresponding maximum a c c e l e r a t i o n s a r e computed i n t h e appendix.
NOZZLF: DESIGN Ektensive e f f o r t w a s expended i n developing t h e nozzle design. To gener- a t e t h e v a r i a b l e - s t a b i l i t y c o n t r o l moments proportional t o inputs and t o meet t h e system requirements, an independent motor-driven a i r r e a c t i o n nozzle was needed. The design c r i t e r i a f o r t h i s nozzle were e s t a b l i s h e d by (1) t h e c h a r a c t e r i s t i c s of t h e bleed air and t h e j e t engine, ( 2 ) t h e requirement t h a t t h e r e a c t i o n f o r c e from each nozzle should be proportional t o t h e command s i g n a l s o as t o adapt t o a c o n t r o l system r e a d i l y , and ( 3 ) t h e requirement t h a t t h e nozzle should y i e l d zero n e t force i n event of l o s s of driving power.
An e l e c t r i c motor-driven nozzle was decided upon s i n c e t h e r e w a s adequate e l e c t r i c a l power a v a i l a b l e on t h e b a s i c a i r c r a f t . Since any weight unbalance f o r a hovering a i r c r a f t should be counterbalanced by weight r a t h e r than by r e a c t i o n force and t h e nozzles a t t h e t a i l require 3 t o 1 counterbalancing t o keep t h e c e n t e r of g r a v i t y above t h e t h r u s t vector, t h e weight of t h e s m a l l motor and gear d r i v e had t o be low. The driving torque w a s t o be kept low enough t o permit use of a low power servomotor.
The concept of t h e o r i g i n a l X-14 p i l o t p i t c h c o n t r o l nozzle, i n which a i r i s discharged from d i a m e t r i c a l l y opposed v a r i a b l e o r i f i c e s , w a s used for a l l s t a b i l i t y augmentation nozzles. Rotary valve motion was used t o c o n t r o l t h e n e t a r e a s i n c e t h e e l e c t r i c motor and gearing adapted e a s i l y t o r o t a r y motion.
Nozzle Torque As a preliminary t e s t , t h e torque needed t o a c t u a t e t h e o r i g i n a l p i l o t ' s p i t c h c o n t r o l nozzle under operating air pressure w a s measured. The r e s u l t s c l e a r l y indicated t h e need f o r a l a r g e reduction i n t h e torque f o r a servomotor drive. There were two components of torque to be overcome; one, the bearing frictional torque caused by unbalanced forces on the nozzle rotor, and the other, the torque caused by the difference in the tangential forces on These torques are illustrated by the cross-sectional the rotor edge areas.
representation of the nozzle (fig. 3 ) .
With the rotor at center, the nozzle pressure inside the rotor acts on the unequal projected rotor areas and so produces a net side force on the rotor bearings. This can be minimized if an additional opening is cut in the rotor, as shown in figure 3, to equalize opposing projected areas. When the rotor is near its extreme position, the internal pressure distribution must balance the net reaction force output of the nozzle and, so again, produces a side force on the rotor bearings. The effects of this force can be mini- mized by antifriction bearings.
The tangential component of rotor torque is due to pressure distribution changes, as the orifices are opened or closed, causing force on the rotor edges, which can produce large torques on the rotor. These, fortunately, tend to center the rotor in most cases.
The rotor edge of the orifice proved to be the most important parameter in meeting the nozzle performance design criteria. The ability of the nozzle to center in a zero net thrust pqsition, when driving power was lost, depended on the choice of edge for the orifice. The torque required to open was dependent on the orifice edge area. For full thrust in one direction, one rotor edge should nominally be covered by the outer shell. However, excessive torque would be required to recover from such a complete closure, so the ori- fice was not allowed to close completely. Several rotor edge modifications were tried (fig. 4) : The 1/8-inch thick square edge orifice had high torque requirements.
A chamfered edge reduced the torque required. An edge was then machined with ribs extending to and reinforcing a thin edge. The results of testing were promising so the ribs themselves were next chamfered to reduce further the rotor edge area.
Rotor edge area control was determined to be the answer to the centering problem by the following experiment. Solder was added to previously tapered edges to see whether it had the opposite effect, and indeed it did. When solder was added to the rotor edge of one orifice to increase the area facing the stream and the rotor edge of the opposite orifice was relieved, the nozzle could be made to stop at any desired position under zero torque. An addi- tional test indicated that the major part of the pressure drop from nozzle pressure to ambient pressure occurs in a very short distance at the orifice edge.
Final Nozzle Design For the final configuration, the rotor was made thinner (1/16 instead of 1/8 inch) to keep the driving torques low andoto enhance the centering characteristics of the rotor edge area. The 120 segment of rotor between o r i f i c e s w a s removed (see f i g . 5 ) and t h e a r e a balancing p o r t i n t h e r o t o r ( s e e f i g , 3) w a s eliminated. The bearing forces caused by t h e unequal oppos- ing projected r o t o r areas were made s m a l l by nearly equalizing t h e pressures i n s i d e and outside t h e r o t o r . This was done by increasing t h e diametrical clearance a s m a l l amount giving a f a i r l y l a r g e circumferential area increase f o r air d i s t r i b u t i o n around both ends of t h e r o t o r .
Data i n f i g u r e 6 f o r t h e f i n a l nozzle configuration show t h e t y p i c a l r o t o r torque required t o obtain a . g i v e n o r i f i c e opening a t t h e operating p r e s - a t t h e s e sure l e v e l i n t h e duct. The servo d r i v e can operate s a t i s f a c t o r i l y torque l e v e l s . For servo a p p l i c a t i o n it i s important t h a t t h e torque be pro- p o r t i o n a l t o r o t o r displacement from c e n t e r p o s i t i o n which, i n t h i s case, means t h a t a r e a and output t h r u s t a r e a l s o proportional t o torque.
Figure 7 i s a photograph of t h e f i n a l nozzle configuration with t h e r o t o r i n i t s centered p o s i t i o n . A l l t h e nozzles had 3 0 ° of a r c width of opening i n They d i f f e r e d i n length of opening t o s a t i s f y t h e a r e a t h e s t a t o r s h e l l .
requirement.
The nozzle r e a c t i o n f o r c e a t t h e temperature and pressure ranges experi- enced on t h e X-14A follows approximately t h e formula (appendix):
F = cvcdf%?1(1.2168 - Pa+&")
Cv = 0.95, discharge c o e f f i c i e n t Using v e l o c i t y c o e f f i c i e n t cd = 0.90, atmospheric pressure P, = 64.7 p s i a , and Patm = 14.7 p s i a , nozzle pressure e x i t a r e a A = 0.80 square inch i n t h e formula gives a m a x i " t h r u s t of 46 pounds f o r t h e roll nozzle.
I n a c t u a l t e s t t h e t h r u s t was 44.8 pounds.
The l i n e a r r e l a t i o n s h i p of r e a c t i o n f o r c e w i t h o r i f i c e a r e a f o r a given duct pressure w a s confirmed by measurement.
Servo Drive The servos p o s i t i o n t h e v a r i a b l e - s t a b i l i t y nozzles i n response t o various e l e c t r i c a l s i g n a l s . Type 0 p o s i t i o n servos (see r e f . 4) were used. For each servo an a l t e r n a t i n g c u r r e n t summing amplifier combines a l l input s i g n a l s and amplifies t h e e r r o r s i g n a l i f t h e nozzle i s not i n i t s commanded p o s i t i o n .
A l l channels have s i m i l a r inputs. For t h e p i t c h channel, t h e inputs a r e : 1. P i t c h r a t e gyro s i g n a l f o r damping 2. Roll r a t e gyro s i g n a l f o r cross coupling i n t o p i t c h a x i s 3. Yaw r a t e gyro s i g n a l f o r cross coupling i n t o p i t c h a x i s 4 . S t i c k p i t c h p o s i t i o n s i g n a l f o r c o n t r o l power P i t c h nozzle p o s i t i o n s i g n a l for p o s i t i o n feedback i n t h e servo 5.
P i t c h nozzle tachometer generator s i g n a l f o r r a t e feedback i n t h e 6.
servo P i t c h summing a m p l i f i e r output s i g n a l f o r gain c o n t r o l 7 .
8. S t i f f n e s s generator s i g n a l f o r s t i f f n e s s c o n t r o l 9. Spare input A magnetic amplifier, accepting t h e output of t h e a l t e r n a t i n g c u r r e n t summing amplifier, d r i v e s t h e two-phase, ten-watt servomotor which operates t h e nozzle. The b a s i c e l e c t r i c a l diagram f o r a s i n g l e a x i s of t h e servo system i s shown i n f i g u r e 8.
The nozzle and d r i v e system were bench t e s t e d . Figure 9 shows t h e nozzle air. The n a t u r a l frequency of t h e unloaded s t e p response with and without nozzle w a s about 9.5 cps, which remained e s s e n t i a l l y unchanged with 75 p s i a a i r i n t h e nozzle. This response introduces l e s s than 20° phase s h i f t f o r normal p i l o t inputs (which a r e l e s s than 2 c p s ) .
The transducers needed t o g e t t h e e l e c t r i c a l s i g n a l s from t h e p i l o t , nozzle, and a i r c r a f t motions included p o s i t i o n transducers,tachometer genera- t o r s , and r a t e gyros. A 5 - v o l t s i g n a l l e v e l w a s s p e c i f i e d t o d r i v e t h e nozzle For t h i s value t h e system w a s not s e n s i t i v e t o e l e c t r i c a l noise f u l l open.
and y e t g r e a t amplification of transducer s i g n a l w a s not necessary.
Rate gyros were used t o produce a s i g n a l proportional t o t h e a i r c r a f t angular v e l o c i t y about each a x i s . The gyros were capable of measuring maxi- mum r a t e s of 6o0/sec i n p i t c h , 60°/sec i n yaw, and 180°/sec i n r o l l . The output voltage of each gyro was amplified and processed t o give a maximum s i g n a l of 5 v o l t s of each phase. P o s i t i o n p i c k o f f s used were microsyns and inductive potentiometers with a l i n e a r range s u f f i c i e n t t o supply t h e neces- s a r y 5 -volt output without amplification.
A t constant a i r c r a f t engine speed t h e engine r o t o r angular momentum i s f i x e d and f o r any a i r c r a f t r o t a t i o n a l r a t e i n p i t c h or yaw a d e f i n i t e amount of torque i s generated by t h e gyroscopic cross coupling. This torque i s pro- p o r t i o n a l t o t h e angular r a t e and i n a d i r e c t i o n dependent upon t h e d i r e c t i o n of a i r c r a f t r o t a t i o n . Since t h e r a t e gyro output i s a l s o p r o p o r t i o n a l t o air- c r a f t r a t e , a n e a r l y exact cancelling of t h e cross-coupling e f f e c t can be achieved i f a proper amount of t h i s s i g n a l i s f e d i n t o t h e proper axis with t h e proper sign.
The s t i f f n e s s generator reduces t h e e f f e c t of short-period unwanted d i s - turbances, such as gusts. This reduction could be obtained with an a t t i t u d e reference as a c o n t r o l system feedback, However, f o r a desired new a t t i t u d e , t h e a t t i t u d e reference would oppose p i l o t control. Thus it was necessary t o allow t h e a t t i t u d e reference s i g n a l (an i n t e g r a t e d r a t e gyro s i g n a l ) t o decay over a period of time. A f i r s t - o r d e r resistance-capacitance l a g c i r c u i t gave an approximation of t h e desired c h a r a c t e r i s t i c s . A time constant of 5 s e c - onds gave adequate s t i f f n e s s b u t r e s t r i c t e d a i r c r a f t maneuvering response.
To c u t out t h e s t i f f n e s s e f f e c t i f a c o n t r o l motion of over 25 percent w a s a r e l a y w a s operated whenever t h e s t i c k p o s i t i o n s i g - commanded by t h e p i l o t , n a l w a s over 1.25 v o l t s (25 percent of maximum).
SYSTEM RESPONSE ON A SIMULATOR In t h e study of a c o n t r o l system which involves new concepts, such as t h e one presented here, it i s advantageous t o study t h e system through simu- l a t i o n using as much a c t u a l f l i g h t hardware i n t h e system as p o s s i b l e . There- f o r e , a low-friction, a i r - b e a r i n g supported, h o r i z o n t a l l y r o t a t i n g beam w a s 10) t o simulate t h e a i r c r a f t c h a r a c t e r i s t i c s f o r one a x i s a t a s e t up ( f i g .
time. With t h i s device it w a s p o s s i b l e t o demonstrate t h e e f f e c t i v e n e s s of t h e c o n t r o l and damping of t h e a i r c r a f t and a l s o t o check t h e operation of t h e e n t i r e system before f l i g h t . I n e r t i a weights and t h e l e v e r arm f o r t h e nozzle were chosen so t h e a c c e l e r a t i o n would be i d e n t i c a l t o a desired a x i s of t h e a i r c r a f t . A i r w a s piped through a r o t a t i n g j o i n t t o t h e t e s t stand and provided both a i r - b e a r i n g pressure and nozzle r e a c t i o n f o r c e s . Nozzle a i r pressure w a s adjusted by a flow valve t o obtain flow r a t e s similar t o those expected i n t h e a i r c r a f t . The servo-drive components were mounted on t h e r o t a t i n g u n i t and connected t o c o n t r o l switches and s i g n a l sources by an overhead cable. This cable w a s brought t o t h e c e n t e r of r o t a t i o n so minimum extraneous torque would be produced on t h e simulator. Without added servo damping t h e r o t a t i o n a l speed of t h e beam decreased from an i n i t i a l 40°/sec t o 10°/sec i n 3-112 minutes. This n a t u r a l damping w a s considered s a t i s f a c t o r y for the simulator t e s t s .
When t h e nozzle servo system w a s used wi'thout damping, t h e angular p o s i - t i o n of t h e beam was r e l a t i v e l y hard t o control. A s servo damping w a s added (by incorporating a s i g n a l from a r a t e gyro) t h e c o n t r o l c h a r a c t e r i s t i c s improved and p o s i t i o n changes could be e a s i l y controlled. When damping w a s reversed i n sign, t h e system soon became uncontrollable.
SAFETY PROVISIONS Safe systems a r e needed f o r any a i r c r a f t b u t they a r e extremely impor- t a n t f o r hovering a i r c r a f t which may have very l i t t l e a l t i t u d e f o r c o r r e c t i v e maneuvers. Electronic f a i l u r e may e i t h e r make t h e servo inoperative or cause it t o drive t o f u l l output. Mechanical f a i l u r e may permit t h e nozzle t o b e f r e e t o ai& center or may cause it t o jam i n an open p o s i t i o n and produce a l a r g e r e a c t i o n force. A switch i s needed for s h u t t i n g o f f t h e power and a valve f o r s h u t t i n g o f f t h e bleed air.
I n order t o implement an e r r o r d e t e c t i o n subsystem f o r each a x i s , a duplicate summing amplifier with a l l nine inputs w a s used. This f a i l u r e d e t e c t i o n a m p l i f i e r w a s designed so t h a t i f it f a i l e d or detected any type f a i l u r e it would operate a cutoff r e l a y , There i s an e r r o r d e t e c t i o n channel f o r each of t h e servo nozzles s o only t h e f a u l t y axis w i l l be c u t o f f , both roll nozzles going o f f if t h e r e i s f a i l u r e i n e i t h e r one.
The p i l o t - o p e r a t e d emergency c o n t r o l s and t h e sequence i n which they a r e as follows. If t h e commands d r i v i n g t h e v a r i a b l e - s t a b i l i t y nozzles used a r e a r e suspect or have caused an unstable f l i g h t condition so t h a t t h e p i l o t wishes t o c e n t e r a l l t h e nozzles, he may t u r n o f f t h e v a r i a b l e - s t a b i l i t y sys- tem switch d i r e c t l y or use a button on t h e s t i c k g r i p which e l e c t r i c a l l y r e l e a s e s t h e same switch. All servos w i l l then d r i v e t o t h e center, or zero force, p o s i t i o n . If one of t h e nozzles i s not driven t o i t s center p o s i t i o n , t h e automatic e r r o r d e t e c t o r w i l l operate and remove e x c i t a t i o n from t h i s noz- z l e . Then, unless it i s stuck, it w i l l a i r c e n t e r . If one of t h e nozzles does not a i r center, a second s t i c k g r i p b u t t o n removes a l l power from t h e v a r i a b l e - s t a b i l i t y nozzles and a l s o energizes t h e motors t o close t h e air valves. A 2-inch valve i s i n each wing duct and a 3-1/2-inch valve i s i n t h e duct t o t h e t a i l . The c l o s i n g time f o r t h e s e valves i s 2 and 4-1/2 seconds, respectively. If a nozzle i s jammed i n an o f f - c e n t e r p o s i t i o n t h e p i l o t must overcome t h e a c c e l e r a t i o n it produces with h i s b a s i c c o n t r o l system u n t i l t h e a i r i s shut o f f . He w i l l then have h i s f u l l b a s i c c o n t r o l for landing. The p i l o t may operate t h e second button first t o t u r n t h e system completely o f f with m i n i " delay.
SYSTEM PERFOWCE I N THE AIRCRAFT Electronic p a r t s of t h e system ( f i g , 11) have given p r a c t i c a l l y no trouble. One f a i l u r e d i d occur i n t h e power supply during a f l i g h t and t h e automatic s a f e t y system operated t o t u r n t h e system o f f . After some f l i g h t experience, t h e automatic s a f e t y system w a s found t o be too s e n s i t i v e , and s i n c e t h e p i l o t s d i d not l i k e t o l o s e t h e v a r i a b l e - s t a b i l i t y c o n t r o l unless it w a s a b s o l u t e l y necessary, t h e s a f e t y system s e n s i t i v i t y w a s reduced by 50 percent. This has worked very well and i s s t i l l much f a s t e r than t h e The motorized nozzles have performed w e l l f o r over l5Ohours p i l o t ' s r e a c t i o n .
The nozzles a r e disassembled p e r i o d i c a l l y f o r inspection and of f l i g h t time.
Several bearings have been replaced as they became preventive maintenance.
noticeably rough and some of t h e roll p i n s holding gears on s h a f t s have loosened, caused backlash, and been replaced. High temperature grease and s i l i c o n e o i l have both been used on t h e bearings, but n e i t h e r seems t o have much advantage on t h e b a s i s of f l i g h t d a t a so far. The nozzle's air centered p o s i t i o n has stayed constant. There has been p r a c t i c a l l y no deformation, p i t t i n g , or d e p o s i t s from t h e bleed air on t h e c r i t i c a l r o t o r edge.
The p i l o t s p r e f e r t h e v a r i a b l e - s t a b i l i t y system t o be i n operation as it makes t h e a i r c r a f t more s t a b l e and e a s i e r t o f l y i n t h e hovering mode. F i r s t , it cancels out t h e unwanted cross coupling and,second, it provides damping and so makes t h e a i r c r a f t respond i n hovering much l i k e it does i n normal f l i g h t . The p i l o t augmented c o n t r o l system permits a yaw r a t e of about 75O/sec and a p i t c h r a t e of about 60°/sec with no noticeable gyroscopic cross coupling between t h e p i t c h and yaw axes. The s t i f f n e s s c o n t r o l augmentation 1 0 .. . . .. .. , .
makes it e a s i e r f o r t h e p i l o t t o keep t h e a i r c r a f t steady under gusty con- d i t i o n s . The s a f e t y f e a t u r e s have provided i n t a n g i b l e b e n e f i t s i n terms of p i l o t confidence.
The system has operated as designed and provided a means of examining t h e c o n t r o l power and damping r e l a t i o n s h i p i n t h e hovering mode for t h e boundaries of good, acceptable, and unacceptable p i l o t control.
Ames Research Center National Aeronautic and Space Administration Moffett Field, C a l i f . , Dee. 4, 1964 APPENDIX FORMULAS AND CALCULATIONS FOR REACTION NOZZLES I n t h i s appendix p e r t i n e n t symbols and formulas are given. Calculations are made t o e s t a b l i s h t h e various nozzle f o r c e s .
SYMBOLS A c r o s s - s e c t i o n a l area, s q ft discharge c o e f f i c i e n t r a t i o , A,/A ' d v e l o c i t y c o e f f i c i e n t cv F r e a c t i o n f o r c e of p i l o t nozzle, l b F' r e a c t i o n f o r c e of v a r i a b l e - s t a b i l i t y nozzle, lb a c c e l e r a t i o n of gravity, 32.2 f t / s e c 2 g I moment of i n e r t i a , slug-ft' L moment arm, f t f n m a s s flow rate, slugs/sec P pressure, p s i a u n i v e r s a l gas constant R T absolute temperature, R flow velocity, f t / s e c
v
G pounds flow r a t e , lb/sec a angular acceleration, rad/sec2 r a t i o of s p e c i f i c h e a t s a t constant p r e s s u r e t o constant volume and Y equal t o 1.4 f o r a i r density, slugs/cu f t P Sub s c r i p t s located upstream located a t o r i f i c e / 3 downstream R0 Air flow
*
p o s i t i o n of sonic v e l o c i t y
-
----
p i t c h P
@--
r roll Sharp edge o r i f i c e t t a i l W wing t i p Y Yaw TOTAL NOZZLE EXIT AREA CALCULSLTED FOR ALL BLEED A I R Mass flow r a t e can be expressed as: and From t h e p e r f e c t gas l a w s it can be shown t h a t : Y+l of 0.90 and knowing t h e r a t i o of Using an experimentally determined cd P,/P1 = 0.5283 a t Mach 1.0 ( s e e ref. 5 ) From engine d a t a $* = 8.5 lb/sec, P 1 = 76 p s i a , and T , = 939' R, S O This is t h e t o t a l allowable solving: A , = 8.5 & / 0 . 4 7 6 7 ( 7 6 ) = 7 .l7 sq i n .
e x i t a r e a t o keep t h e bleed flow desired.
TOTAL REACTION FORCE AVAILABLE FROM ALL NOZZLES The t o t a l r e a c t i o n force, i d e a l l y , can be expressed as:
F = m , V , + A,(P, - P3)
Because of f r i c t i o n a t t h e o r i f i c e edge, w i l l not be obtained over t h e V * A,, b u t w i l l be l e s s a t t h e outer edges of t h e flow stream. This, whole a r e a i n e f f e c t , reduces t h e sonic t h r o a t a r e a t o give an average a r e a of sonic flow of “A+, s o a c t u a l l y , Using t h e r e l a t i o n s h i p s and P, = p,RT, and I& = p,A,V,CV, ( V + ) 2 = yRT, keeping sonic flow:
F = CvCdA2Pl ( 1 . 2 6 8 - 3
With Cv = 0.95 from t e s t d a t a and P3 as atmospheric pressure,
F = (0.95)(0.90)(7.17)(76)[1.268 - (14.7/76)] = 500.7 pounds t o t a l r e a c t i o n
f o r c e .
DIVISION O F TOTAL REACTION FORCE AMONG THE NOZZLES Ia
For any nozzle: F = -
L I~ 0.90 ap
IPOGP
Fp = FP = L t L t F - Iyo5r 9 = 0.284 a, y - Lw Irar
ar = 1.0 +
Fr = L , The sum of t h e separate individual forces must equal t h e t o t a l available; a r e not independent, and s i n c e Fr i s t h e l a r g e r , it must be used. Fr and Fy f F’ f Fr f Fi + F $ F F t o t a l = p p I ar = -- 500*7 - 1.84 rad/sec2 272.2
4 = 0 . 9 % = 1.66 rad/sec2
ar = 1.84 rad/sec2
4 = 0.9 9 = 0.83 rad/sec 2
% = 0.5 % = 0.92 rad/sec2
4 = 0.9 "y = 0.47 rad/sec2
% = 0.284 % = 0.52 rad/sec2 F = 9.8 lb P F ' = 88.0 lb P Fr = 127.2 lb Non-independent Fy = Fr cos 4 5 ' F~ = 127.2 (0.707) Fi = 114.5 lb Fi = 73.2 lb Fy = 90.2 lb 500.7 1b RFFEmNCES 1. Rolls, L. Stewart,and Drinkwater, Fred J., 111: A F l i g h t Determination of t h e Attitude Control Power and Damping Requirements f o r a Visual Hovering Task i n t h e Variable S t a b i l i t y and Control X-14A Research Vehicle. N A S A 'I" D-1328, 1962.
2 .
O'Malley, James A., Jr., and Landphair, Lee C . : The X - 1 4 VTOL A i q l a n e - A Design Tool. SAE Paper 95B, 1958.
3. Cooper, George E.: Understanding and I n t e r p r e t i n g P i l o t Opinion. Aero.
Eng. Rev., vol. 16, no. 3, March 1957, pp. 47-51, 56.
4. Savant, C. J.: Basic Feedback Control System Design. McGraw-Hill, New York, 1958.
5. Ames Research S t a f f : Equations, Tables, and Charts for Compressible Flow.
NACA Rept . 1135, 195 3.
A-2 84 73 - 9 Figure 1.- Photograph of X-14A hovering.
I -
VARIABLE STABILITY
VARIABLE STABILITY-
,/ PITCH 88.0 I b
PILOT COMMAND
E LECT R I CA L LY 0 PERAT E D
IR SHUTOFF VALVES
AIR SHUTOFF VALVE
ENGINE B L E E D / L
VARIABLE STABILITY
A I R CONNECTIONS
ROLL ,T 57.25 I b
Figure 2.- Nozzle l o c a t i o n md forces on X-14A aircraft.
X ATMOSPHERIC PRESSURE REACTION FORCE/ / ROTOR EDGE FORCE\ I / / REACT1ON FORCE LATER R
PROJECTED I-. ---
AREAS NET FORCE ON ROTOR BEARING DUE TO UNEQUAL PROJECTED AREAS NOZZLE PRESSURE CUTOUT I N ' Y ROTOR TO
i NET FOI
RCE EQUALIZE ROTOR ON BEA RING PROJECTIONS EQUAL ' TO NET REACTION FORCE
(a) ROTOR AT CENTER
( b ) ROTOR NEAR ONE EXTREME
Figure 3 . - Forces acting on a prototype nozzle.
SQUARE
CHAMFER
RIBBED
CHAMFERED
R I B S
Figure 4 . - Nozzle r o t o r edge modifications.
A-31765 Figure 5 . - F i n a l nozzle rotor.
-
60 -
- --
--
40 -
I - --
z
--
- a 20- SSURE
W
-
--
--
--
--
S O Y I I I I 3- I I I I I
io0 80 60 40 20 o 20 40 60 ao 1 0 0
NOZZLE OPENING I percent of maximum
Figure 6.- Typical nozzle torque characteristics.
A-31764 Figure 7.- F i n a l nozzle assembly.
rc w
ERROR DETECTOR
AND AUTOMATIC
SHUTOFF
h I I
SPARE l N P U T d \ I
PILOT COMMAND
NOZZLE
SIGNAL - y\
- \
ROLL, PITCH NOZZLE
AND YAW RATE
DRIVE
G Y R O SIGNALS
FEEDBACK
1 POSITION I
I 1 I I
STIFFNESS
I I FEEDBACK I
GE N E R AT0 R W IT H I
I I I
I MANEUVERING I I i
CUTOUT
FEEDBACK
Figure 8.- X-14A control system diagram for one axis.
+ 50 -
t
c
-
Q)
G o
Q
c - 4 k.1 sec
e 9
2 - 5 0 -
-
z ( a ) NO A I R P R E S S U R E
k! +50 -
-
W
A
N 0
N
0 -
z
-50 -
( b ) 75 pSia NOZZLE AIR PRESSURE
Figure 9 . - Nozzle step response.
A-26690 Figure 10.- Single-axis aircraft simulator.
A-32540-6 Figure 11.- Mounting of electronic units.
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