Application of the model technique to a variable-stability helicopter for simulation of vtol handling qualities

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(CATEGORY) ‘coDE’os Microfiche (MF) ’ (NASA CR OR TMX OR AD NUMBER) ff 653 July65 APPLICATION OF THE MODEL TECHNIQUE TO A VARIABLE-STABILITY HELICOPTER FOR SIMULATION O F VTOL HANDLING QUALITIES By John F. Garren, Jr., and James R. Kelly NASA Langley Research Center Langley Station, Hampton, Va., U.S. A.

Presented at the 27th Meeting of the AGARD Flight Mechanics Panel Rome, Italy October 11-12, 1965

NATIONAL AERONAUTICS AND

SPACE ADMINISTRATION

WASHINGTON APPLICATION OF THE MODEZ TECHNIQUE TO A VARIABLE-STABILITY HELICOFTE3 FOR SIMULATION OF VTOL HANDLING QUALITIES

By John F. Garren, Jr. ,* and James R . Kelly*

N A S A Langley Research Center SUMMARY 1349y4 I n order t o provide a means f o r accurate in-flight simulation of V/STOL a i r c r a f t , a computer model technique has been adapted t o a variable-stability helicopter. Unlike t h e stability-derivative simulation technique, which i s usually employed i n variable-stability a i r c r a f t , the model approach produces a response which is essentially independent of the dynamics of the test vehi- cle. The a i r c r a f t response, therefore, is a function only of the evaluation p i l o t ' s control inputs and the dynamics which are programed i n t o the analog computing equipment.

In-flight t i m e h i s t o r i e s of the helicopter response and the corresponding commanded response are presented t o i l l u s t r a t e the effectiveness of t h e tech- nique. The results indicate t h a t the model technique does, i n fact, provide a feasible, accurate, and flexible approach t o in-flight simulation.

INTRODUCTION The c r i t i c a l need f o r improving the v a l i d i t y and coverage of existing VTOL handling-qualities c r i t e r i a is well recognized. In the absence of V T O L a i r c r a f t of handling-qualities suitable f o r conducting the required studies, the bulk m u s t come from simulation. It is essential, moreover, t h a t the particu- data larly c r i t i c a l areas be explored by means of airborne simulation because of intangible influences of the p i l o t environment and f l i g h t task. I n the past, however, airborne simulation has been hampered by an i n a b i l i t y t o represent a wide range of characteristics with a sufficient degree of accuracy. The inac- curacies have stemmed from a lack of information r e l a t i v e t o the characteristics of t h e basic vehicle, as w e l l as from the complexity involved i n a l t e r i n g some of the more important characteristics as discussed i n reference 1.

The miniaturization of analog computing equipment during recent years has made it possible t o circumvent these problems by applying ground-based s i m u l a - When applied t o airborne t i o n techniques t o variable-stability a i r c r a f t .

simulation t h i s method i s commonly referred t o as the model simulation tech- nique. I n t h i s technique the equations of motion, which represent the a i r c r a f t

*

Aerospace Technologist.

L-4499 In characteristics being simulated, are programed into an onboard computer.

the computer, the desired response is generated in accordance with the $0- gramed equations on the basis of control inputs and motion-sensor feedbacks.

Differences in the desired response and the vehicle's actual response are used to form an error signal which drives the control surfaces so as to eliminate the discrepancy.

The modification of a prototype helicopter for adaptation of the computer- Following devel- model simulation technique w a s initiated by the NASA in 1961.

opment and documentation of the simulation capability in 1962, actual research flights began in 1963. The purpose of this paper is to present a description of this technique as it has been applied to low-speed VTOL research. Limita- tions of this technique encountered under operating conditions are discussed.

In-flight time histories are presented to illustrate the effectiveness of the model technique.

A general description of the variable-stability helicopter is also included.

LIST OF S Y M B O L S forces along the respective body axes, lb moments about the body X, Y, and Z axes, respectively, lb-ft moments of inertia about the respective body axes, slug-ft2 mass of aircraft, slug used as subscript to designate commanded response (i.e., the computer response) used as subscript to designate actual response (i.e., the response of the test helicopter) control deflection about the three body axes indicated and along k,6y,6z,6p the body Z axis, respectively, in.

P rolling angular velocity, rad/sec 9 pitching angular velocity, rad/sec r yawing angular velocity, rad/sec U forward component of velocity, ft/sec V side component of velocity, ft/sec W n o m component of velocity, ft/sec a angle of attack, rad components of simulated gust velocity, f t / s e c 71,72,73,74 z absolute altitude, f t Unless otherwise indicated, when any of the defined symbols are used as subscripts, the derivative w i t h respect t o that parameter is indicated. Dots over a symbol indicate a time derivative w i t h respect t o t h a t parameter.

GENERAL S m I O N CONSIDERATIONS Equations of Motion The terms which are of primary interest because of t h e i r first-order e f f e c t s on the response of VTOL a i r c r a f t are shown below i n the form of linear, quasi-static equations of motion:

i = u r + f(q) + YV mV

( 5 ) The coefficients i n the equations need not be constant since nonlinear elements are available i n the computing equipment t o permit variation i n the coefficients w i t h airspeed and a l t i t u d e o r other parameters. Similarly, varied relation- ships between control position and the corresponding accelerating moment can be handled.

Because i n the a i r c r a f t there are o n l y four independent sources f o r pro- ducing moments and forces (moments about each of the three body axes and a l i f t force o r t h r u s t along the body -2 axis), it i s possible t o alter the basic vehicle response f o r only the degrees of freedom expressed by the first four equations.

In the absence of independent sources f o r producing forces .along, the body -X and -Y axes (which would necessitate i n s t a l l a t i o n of additional propulsion systems) it is necessary t o l i v e with the characteristics of the basic aircraft f o r the latter two degrees of freedom, equations ( 5 ) and (6).

It may be seen from inspection of the l a t t e r two equations, however, t h a t only the last term of each, the drag, is a function of the particular a i r c r a f t ; the other terms are independent of a i r c r a f t configuration so t h a t the basic vehicle i s otherwise inherently correct. Fortunately, therefore, the i n a b i l i t y t o alter the response f o r these two degrees of freedom represents only a minor limita- tion, particularly at low speeds where drag e f f e c t s a r e usually quite small.

Fromthe p i l o t viewpoint,' not duplicating the drag term r e s u l t s i n a devia- t i o n fromthe correct relationship between a i r c r a f t a t t i t u d e and the steady- s t a t e linear velocity. S t r i c t l y speaking, there can also e x i s t s l i g h t differ- ences i n the long-term response following control inputs, but such differences are so small as not t o be perceived by the p i l o t i n the majority of cases. For example, if one wished t o simulate a t i l t - w i n g VTOL through the conversion from hover t o cruise f l i g h t , the fuselage a t t i t u d e would not be correctly duplicated even though the dynamics and response t o control inputs would be e s s e n t i a l l y correct. From a handling-qualities standpoint such e f f e c t s are probably minor i n comparison w i t h the other parameters which are being studied and, therefore, do not currently j u s t i f y the increase i n complexity which would be associated w i t h adding sideward and forward (and rearward) thrusting engines.

Mechanization of Equations The solutions f o r the first three equations of motion as l i s t e d previously are desired i n terms of angular velocities about the respective axes; from the fourth equation, the normal acceleration is required. These solutions are obtained i n real time and are the responses which are used t o connuand t h e vehi- cle motion. I n order t o obtain the solution f o r the first three equations, the outputs f r o m the p i l o t controls and from various motion sensors are summed i n accordance with the specified equations of motions t o produce a voltage propor- t i o n a l t o the desired angular acceleration which i s 'integrated i n turn, t o yield a signal proportional t o the desired angular velocity. For i l l u s t r a t i v e pur- poses the mechanization of the lateral-directional equations of motion, equa- t i o n s ( 2 ) and (3), i s shown schematically i n figure l. It should be noted t h a t the coefficient values, which are s e t on the potentiometers, have been normal- ized with respect t o a i r c r a f t i n e r t i a s so t h a t the e f f e c t of i n e r t i a on the periods and time constants is automatically taken i n t o account. The method of mechanization shown i n the figure a l s o accounts f o r the proper degree of coupling, or interaction, between the axes.

A further example of the mechanization which i s required is the switch-over from air-referenced t o ground-referenced signals at speeds below about 30 K where some sensors, such as the s i d e s l i p vane, become unreliable. The switch t o pro- vide t h i s function is shown i n the schematic diagram. It i s pointed out, more- over, t h a t for speeds above 30 K, the s i g n a l proportional t o the a i r c r a f t side- ward velocity v is generated by passing the s i g n a l from the angle-of-sideslip vane through a potentiometer which is set for the intended forward speed. For test conditions where the speed will not be held constant, the potentiometer cai be replaced by an element which accounts for changes in velocity. Varia- tion of the coefficients with speed can be handled in much the same manner.

DESCRIPTION OF EQUIPMENT Test Vehicle A photograph of the NASA variable-stability helicopter is shown in fig- ure 2 . The gross weight of this vehicle is 15,500 pounds and its maximum speed is 140 knots. The vehicle has been demonstrated to rearward and sideward velocities up to 25 knots and to a normal acceleration of l.5g. It should be noted that this aircraft is a prototype and these restrictions are not typical of production models. A more complete documentation of the characteristics of the basic aircraft is given in reference 2.

As in the case of any simulator, the maximum accelerations and velocities which can be simulated are limited to the corresponding capabilities of the test vehicle. These limitations for the angular degrees of freedom for the NASA test helicopter are given below.

Maximum angular Maximum angular Axis acceleration, velocity, rad/sec2 rad/sec Pitch 1.7 "he angular acceleration in yaw is considered marginal although its near-zero angular velocity damping permits extremely high angular velocities to be developed. At any rate, for the simulation of dynamics pertinent to aircraft as large as, or larger than, the test helicopter, these acceleration and rate Difficulties are sometimes capabilities have generally proven to be adequate.

encountered, however, in the simulation of higher response associated with small aircraft.

Variable-Stability System The variable-stability system (ref. 3) installed in the helicopter is com- posed of three major components; namely, a modified control system, an analog The location of each comgonent in the overall computer, and a sensor group.

The block in the figure labeled "signal system is illustrated in figure 3.

plugboard" provides the interface between the various components. Although the function of the latter two components has been discussed previously, additional information is included i n t h i s section.

Modified control system.- The p i l o t controls on the right-hand side of the cockpit, consisting of the conventional center stick, rudder pedals, and collective s t i c k were modified t o a "fly-by-wire" system; that is, the motion of these controls produces only e l e c t r i c a l voltages which can, i n turn, be used the control surfaces. The left-hand controls a r e unmodified and are t o drive continuously monitored by the safety p i l o t whose duty it i s t o take over i n the Further modification included pro- event of a malfunction or other emergency.

vision f o r the conversion of e l e c t r i c a l voltages t o control surface displace-

ments. There a r e four separate but i d e n t i c a l Channels i n the system - one each

f o r the pitch, r o l l , yaw, and v e r t i c a l degrees of freedom.

Since f a i l - s a f e features were not designed into the various components Of the variable-stability system, several safety provisions were incorporated i n the modified control system. These provisions included a control-limiting system, a safety p i l o t override feature, and several disengage modes. The control-limiting system permits a variation i n the authority of the variable- 10 percent t o 100 percent. For example, when t h e s t a b i l i t y system from authority is s e t a t 10 percent, the system i s capable of commanding only 10 per- cent of the t o t a l control surface travel. Although the simulations a r e normally run at 100 percent authority, the i n i t i a l engagement on each flight i s made a t a reduced authority (about 30 percent). The disengage modes, which revert con- t r o l of the a i r c r a f t t o the safety p i l o t i n the event of an emergency, include e l e c t r i c a l disengage buttons on each of t h e p i l o t s ' controls a s well as a As a precaution against redundant mechanical switch on the instrument panel.

hardover failures, the override feature allows the safety p i l o t t o overpower commands by the variable-stability system without having t o f i r s t disengage the system.

Analog computer.- The computing equipment which i s located i n the cabin i s "his computing equipment shown i n figure 4 along with the signal plugboard.

consists of two off-the-shelf computers, which are slaved so t h a t both may be operated from a single control panel. The equipment, as shown, i s sufficient f o r programing three degrees of freedom, but i s being expanded t o handle t h e fourth. The computing elements currently i n s t a l l e d i n the NASA helicopter include forty (40) operational amplifiers, sixteen (16) integrators, forty- eight (48) potentiometers, and twenty-four (24) nonlinear components. I n addi- t i o n t o computing the command response, several complementary functions are performed. Some elements are used i n establishing quasi-static conditions a t as canceling, the instant of engagement. This function, commonly referred t o s t a r t s the outputs from a l l the sensors at zero, which also prevents t r a n s i e n t s the system. S t i l l another f'unction served by t h e computer upon engagement of i s the correction of various motion sensors f o r t h e i r location r e l a t i v e t o the a i r c r a f t center of gravity.

Sensors.- Insofar a s possible, the locations of the various sensors were chosen with regard t o t h e i r respective function. "he angular velocity and angular acceleration sensors may be mounted i n any convenient location so that t h e i r position w a s selected on the b a s i s of minimum vibration. Most other sensors, on the other hand, a r e sensitive t o location with respect t o the air- c r a f t center of gravityj and, where a center-of-gravity location i s not feasi- ble, corrections t o t h e i r outputs m u s t be made. For example, t h e angle-of- attack sensor must be mounted ahead of t h e a i r c r a f t t o minimize rotor downwash effects. I n t h i s position, however, the vane i s sensitive t o pitching angular velocity. By properly summing the output f r o m the vane w i t h the output from the longitudinal r a t e gyro, however, the true angle of a t t a c k i s obtained.

Similarly, corrections a r e required f o r the angle-of-sideslip vane and f o r the l i n e a r accelerometers. I n addition t o contributing t o the solution of the command response, the sensor outputs a r e recorded f o r correlation with the p i l o t ratings and comments, and, i n some cases, they a r e used t o actuate cock- p i t displays.

RESPONSE COMMAND MJITHOD The method used f o r generating a signal proportional t o the desired, o r commElnded response was discussed i n a preceding section. This section describes the technique which forces the vehicle t o obey the commanded response. The basic command technique, a s i l l u s t r a t e d i n figure 5 , i s nothing more than a closed-loop servomechanism employing a r a t e error signal. I n t h i s case the computer p r w i d e s the c o m n d e d angular r a t e qc and the feedback, or actual r a t e , i s provided by a r a t e gyro which i s mounted i n the a i r c r a f t .

qh Limitation of the Basic Command Method EPfect of time delays.- Inspection of simultaneous time h i s t o r i e s of the commanded and the actual response f o r systems of the type shown i n figure 5 generally indicate a time lag, o r phase shift between the two, even though the t i m e h i s t o r i e s might appear essentially identical i n other respects. Any such delays are, of course, undesirable and i f they appraoch 0.2 second o r so, they tend t o become discernible t o the p i l o t a s a delay i n t h e response t o control inputs. A more subtle effect of the time delay than the mere phase shift between t h e commanded and t h e actual response i s the e r r o r which it produces i n the commanded response i t s e l f .

There a r e then, i n f a c t , three d i s t i n c t responses which should i d e a l l y be

i d e n t i c a l - the theoretical response, the commanded response, and t h e actual

response. Although one i s tempted t o assume that the commanded response is automatically the same as t h e t h e o r e t i c a l response, this i s s t r i c t l y true only f o r simple dynamics (for example, zero-order and first-order responses) where t h e solution of the equations of motion does not involve the motion sensors.

A zero-order response results from an acceleration system i n which case the only input t o t h e summing amplifier i n figure l w o u l d be the output f r o m t h e p i l o t control. Similarly, no motion sensors a r e used i n computing a first-order response, a r a t e system, i n which case there would be two inputs t o the summing

amplifier - the p i l o t control and a r a t e feedback f r o m t h e computed angular

velocity. For most other instances the motion sensors a r e required f o r Solu- t i o n of t h e commanded response and time delays inherent i n the basic command method are fed back through the sensors which cause the commanded response t o deviate fromthe theoretical response.

"his, i n turn, causes the actupl respony;e, which closely duplicates t h e commanded response, t o be i n error. The error, thence, tends t o be self-generating, so t h a t t o accurately duplicate the theoretical response f o r long intervals of time, say 30 seconds, would require t h a t the vehicle-following time delay not exceed a few hundredths of a second.

I n general, high-frequency responses having low damping r a t i o s a r e the most adversely affected by the time delay.

For t h e evaluation of handling q u a l i t i e s during precision tasks, such as instrument approaches and hovering over a spot i n turbulence where the p i l o t control frequency i s on the order of one input per second, exact duplication of the theoretical response f o r long periods of time i s not mandatory. I n such cases, where the p i l o t may be considered as an active element i n the control loop, the i n t e r v a l of prime importance, a s shown by several handling-qualities investigations during recent years, i s the first 2 o r 3 seconds of the response following t h e control input. It does not appear unreasonable, however, t o require t h a t the system have an accuracy of no l e s s than 80 percent during the first 10 seconds following a disturbance, e i t h e r by the p i l o t o r by the simu- lated turbulence.

Source of time delays.- Since the reduction of time delays i s the key t o achieving an accurate response, it is necessary t o understand t h e i r source.

One source i s the dynamics of the basic test vehicle which typically contributes time lags on the order of 0.1 second. The second source i s d i r e c t l y dependent on the error-signal gain (amplification of the e r r o r signal) which can be attained. "he error-signal gain i s defined here as the angular acceleration which the helicopter develops t o cancel a u n i t e r r o r i n angular velocity and i s

" 1

defined as: G = . It i s noted, therefore, t h a t G has units of second

% - qh

and, i n the absence of control system time delays, can be considered a s approx- imately representing the reciprocal of the time required f o r the vehicle t o achieve 63 percent of any commanded rate. For example, assuming a s t a t i c gain of lO/sec f o r G, the actual response w i l l l a g the commanded response by 0.1 second. It should be pointed out that the time delays from each of the sources mentioned a r e not d i r e c t l y additive since the e r r o r signal overcontrols i n an e f f o r t t o reduce the basic vehicle t i m e delay.

Since t h e time delay associated with the closed-loop dynamics i s the reciprocal of the e r r o r signal gain, it would be desirable t o a t t a i n an i n f i n i t e gain.

A s i n t h e case of any p r a c t i c a l system, t h e characteristics of the var- ious loop-elements l i m i t the m a x i m u m allowable gain t o some f i n i t e value, beyond which t h e control system w i l l l i m i t cycle, i.e., a self-sustained oscil- l a t i o n of high frequency and constant amplitude w i l l e x i s t i n t h e control loop.

Gains on the e r r o r signal which were attainable i n flight using the basic sys- tem of figure 5 resulted i n time delays, the worst of which w a s about 0.3 sec- ond while the best w a s somewhat less than 0.1 second. Analytical studies based on t h i s worst case indicated t h a t a wide range of damping r a t i o values, including a damping r a t i o of zero, could be s a t i s f a c t o r i l y simulated only f o r periods greater than 7 seconds.

Even a t low speeds it i s possible f o r Small a i r c r a f t , i.e., a i r c r a f t with low moments of i n e r t i a , t o exhibit periods

a

somewhat shorter than t h i s . It was apparent therefore t h a t some modification t o .the ba'sic technique was necessary.

Modification of Basic Command Technique

- Lead.- Aside from the long-term inaccuracies which tend t o accumulate

when there i s an inadequate gain on t h e r a t e error signal, the l a g produced i n the i n i t i a l or transient response following control inputs i s no doubt the most adverse e f f e c t from a handling-qualities standpoint. Time delays on the order of 0.3 second a r e within the p i l o t capability of observation and would there- fore r e s u l t i n pessimistic p i l o t ratings for the simulated characteristics.

I n order t o overcome t h i s delay, a lead network, a s shown i n figure 6, was This input i s scaled so as t o produce the cor- added t o the basic technique.

r e c t i n i t i a l acceleration following motion of the control.

Additional lead can be provided t o further reduce the l a g i n the actual response by feeding the motion sensor outputs i n t o the control system through the lead network. The inputs from the sensors are scaled so t h a t the s t a b i l i t y derivatives of the basic t e s t vehicle a r e a r t i f i c a l l y a l t e r e d t o match t h e s t a b i l i t y derivatives being simulated. Such use of lead i s the sole simulation method used i n conventional airborne simulators, so t h a t the present simulation technique i s i n a c t u a l i t y a hybrid of the pure model technique and the con- ventional o r s t a b i l i t y derivative technique.

Inasmuch a s t h e use of lead i n t h i s application, however, i s only a second-order refinement, t h e matching o f the basic vehicle characteristics t o the desired characteristics need be only approximate t o yield the desired r e s u l t . Assume, f o r example, t h a t t h e characteristics of the basic t e s t vehi- c l e a r e only approximately known so t h a t the best simulation which can be achieved using only lead i s about 60 percent. Assume, further, t h a t the basic, o r unmodified model simulation technique i s capable o f compensating f o r 80 percent of the difference between the desired response and t h e inherent response of t h e basic t e s t vehicle f o r some range of characteristics. By com- bining t h e two methods the expected accuracy can be figured approximately a s 60 percent plus 0.8 (4.0 percent), o r 92 percent. I n other words, even by adding i s uncertain by a margin of 40 percent, the error i n the compensation t h a t It is seen, therefore, simulation i s reduced from 20 percent t o 8 percent.

t h a t t h e lead need not be precise t o be effective.

Integrator loop.- There a r e many characteristics i n t h e basic test vehi- c l e which contribute minor e r r o r s t o t h e simulated response, but which cannot be compensated f o r by using lead ( p a r t l y because of a lack of appropriate sensors and p a r t l y because of the added complexity). T r i m changes of the basic vehicle and i n e r t i a coupling effects a r e representative examples of such char- a c t e r i s t i c s . An effective compensation network f o r t h i s purpose, which was suggested by National Research Council personnel who a l s o use it i n t h e i r v a r i a b l e - s t a b i l i t y helicopter, i s the integrator network shown i n figure 6.

A s t h e name implies, t h i s network integrates any e r r o r i n t h e angular velocity and therewith feeds i n additional control.

A s i n t h e case of t h e gain on t h e rate e r r o r signal, only a limited gain can be t o l e r a t e d on t h e integrated-rate e r r o r signal. I n f a c t , t4e addi- t i o n of the integrated signal represents a compromise i n t h a t it becomes necl essary t o reduce t h e rate-error-gain since t h e overall e r r o r gain i s r e l a t e d t o t h e square root of the sum of the squares of these individual gains. Fur- thermore, the overall gain which can be achieved i s somewhat lessened because of t h e phase characteristics introduced by t h e i n t e g r a t o r loop. Addition of t h e integrator loop provides, nonetheless, a net improvement i n t h e overall response, p a r t i c u l a r l y f o r t h e low-frequency response. A l s o , with regard t o elimination of external disturbances, t h e integrator provides a long-term a t t i - tude memory whereas the pure r a t e e r r o r signal provides only high damping.

I n general, then, the lead improves t h e high-frequency response; t h e integrator loop, the low-frequency response; and the basic r a t e e r r o r signal operates over t h e e n t i r e spectrum.

Results Using Modified Techniques Figure 7 i s a f l i g h t t i m e history obtained using t h e modified technique.

The t r a c e labeled "qc" i s t h e commanded pitching angular velocity and t h e one "qhf' i s t h e a c t u a l angular velocity. Both t r a c e s a r e recorded on labeled approximately t h e same gain. The timing marks, t h e v e r t i c a l l i n e s , a r e a t 1-second intervals. An estimate of t i m e delay i n t h e vehicle-following may be A c a r e f u l inspec- obtained by comparing the t i m e between corresponding peaks.

t i o n of t h i s figure reveals t h a t t h e o v e r a l l t i m e delay does not exceed 0.1 second.

A s discussed i n an e a r l i e r section, t h e basic technique w a s substantially limited i n t h e range of c h a r a c t e r i s t i c s which could be accurately simulated of an appreciable time delay r e s u l t i n g from limitations on t h e e r r o r because signal gain. The modified technique, on the other hand, provides a capability f o r simulating o s c i l l a t o r y responses with periods as short as 4 seconds a t very The low damping r a t i o s , and even shorter periods a t higher damping r a t i o s .

primary obstacle t o f u r t h e r increasing t h e range of response which can be s i m - ulated i s the t i m e delay associated with t h e basic vehicle. The present capability, however, has thus f a r proven adequate f o r t h e simulation of char- a c t e r i s t i c s pertinent t o low-speed handling q u a l i t i e s f o r VTOL a i r c r a f t .

An additional c r i t e r i o n against which t h e model simulation technique was judged w a s i t s a b i l i t y t o eliminate external, o r unprogramed, disturbances such as t r i m changes of t h e basic test helicopter. As i s t h e case with other tandem-rotor helicopters, t h e t e s t helicopter exhibits a strong t a i l - s i t t i n g tendency during decelerating f l a r e s , which must be countered by forward motion of t h e longitudinal control. I n order t o t e s t t h e system capability i n t h i s flight records of t h e evaluation p i l o t ' s control and of t h e helicopter respect, control surfaces were obtained as t h e evaluation p i l o t f l e w t h e a i r c r a f t through the c r i t i c a l maneuver. These control-motion t i m e h i s t o r i e s are com- 8. This figure i l l u s t r a t e s c l e a r l y that t h e evaluation p i l o t ' s pared i n figure control remained trimmed near zero throughout t h e maneuver, despite t h e f a c t 25 percent of i t s t h a t t h e control system of the basic helicopter moved about Total t r a v e l i s used here a s the dis- t o t a l t r a v e l t o fight the t r i m change.

tance beOween the control t r a v e l l i m i t s . It was concluded from such t e s t s t h a t the model technique effectively eliminates any unprogramed disturbances.

APPLICATION OF T H E MODEL SIMULATION TECHNIQUE A s implied i n the introduction, the principal use of the N A S A variable- s t a b i l i t y helicopter i s i n the development of general c r i t e r i a f o r VTOL handling q u a l i t i e s . For t h i s purpose, rather than simulating the detailed characteristics of specific configurations, a wide range of different param- e t e r s a r e systematically evaluated f o r a variety of tasks. The merits of the model simulation technique f o r such studies a r e best emphasized by a discus- sion of a few of the recent contributions t o handling q u a l i t i e s which it has made possible.

The problem of applying helicopter c r i t e r i a t o V T O L a i r c r a f t i s particu- l a r l y c r i t i c a l i n the specification of control power, which, though r e l a t i v e l y inexpensive i n helicopters, must be provided a t the d i r e c t expense of i n s t a l l e d power i n many VTOL configurations. It i s thus important t h a t minimum require- n;er,ts be accurately determined. Since one of the fundamental requirements f o r control power i s maneuvering, an extensive investigation ( r e f . 4 ) of maneu- vering requirements w a s conducted during which t r i m changes and disturbances were eliminated by use of the model technique. Although differences i n t r i m change characteristics and gust susceptibility have been the principal c r i t i - cism of applying helicopter experience t o VTOL a i r c r a f t , t h e r e s u l t s of t h i s study showed close agreement with AGARD Report 408 ( r e f . 5 ) which was based largely on helicopter experience. The maxim difference f o r any of the axes was only 20 percent; the general v a l i d i t y of t h i s portion of current c r i t e r i a i s thus more firmly established.

Precision tasks, such a s hovering over a spot and square hovering pat- terns, were a l s o performed during the control power investigation and it was soon apparent t h a t wide variations i n either control power or s e n s i t i v i t y (angular acceleration per u n i t control) had l i t t l e , i f any, e f f e c t on the p i l o t ratings. I n t h e absence of disturbances, the visual precision tasks became t r i v i a l even f o r low values of damping, and the a i r c r a f t could be "balanced" with very l i t t l e p i l o t e f f o r t . The overwhelming e f f e c t of dis- turbances on the precision hovering t a s k is clearly evident by comparison of studies where disturbances, could not be elimi- these r e s u l t s with previous nated. The a b i l i t y t o i s o l a t e the various parameters and t o examine t h e i r e f f e c t s individually has contributed t o a clearer understanding of the overall handling-qualities picture.

The e f f e c t s of t r i m change, s t a t i c s t a b i l i t y , and simulated turbulence on these r e s u l t s will be the subject of f'uture investigations. I n t h e case of t h e yaw axis, the e f f e c t s of s t a t i c directional s t a b i l i t y have already been examined with t h i s equipment and a r e reported i n reference 6.

A f l i g h t investigation conducted with an on-off type control using t h i s equipment provides another example of t h e p o t e n t i a l offered by t h e model tecp- nique. This investigation w a s conducted t o determine t h e f e a s i b i l i t y of using an on-off control system (as opposed t o t h e conventional proportional system) f o r V/STOL operation. The parameters which were investigated included t h e s i z e of the control deadband, control power e f f e c t s , and angular velocity damping effects. During these tests extremely low control powers were inves- t i g a t e d and were found t o give satisfactory maneuver capability. I n f a c t , satisfactory control f o r maneuvering w a s obtained a t about one-quarter t h e control power needed f o r proportional control systems. Total moments equiva- l e n t t o l e s s than 3 percent of t h e t o t a l available control-surface t r a v e l of t h e basic a i r c r a f t were explored (values t h i s low were not satisfactory, how- Since t h e basic a i r c r a f t exhibits t r i m changes on t h e order of 25 per- ever).

cent of its t o t a l control-surface displacement, as discussed previously, this study could not have been accomplished had not t h e simulation technique been capable of eliminating t r i m changes.

Although t h e equipment, as i n s t a l l e d i n t h i s a i r c r a f t , i s best s u i t e d f o r establishing general handling-qualities c r i t e r i a , such an a i r c r a f t can be used effectively as a t r a i n i n g device t o b e t t e r acquaint t e s t p i l o t s with t h e e f f e c t s of various s t a b i l i t y derivatives on f l y i n g q u a l i t i e s . Also, by s i m - u l a t i n g a i r c r a f t t h a t are s t i l l i n the design stage, it often i s possible t o detect potential problem areas and t o determine t h e direction and magnitude of changes t o correct t h e deficiency. A s s t i l l another application, it i s always advantageous t o simulate the c h a r a c t e r i s t i c s of newly b u i l t and yet unflown a i r c r a f t so t h a t t h e p i l o t can gain experience with new o r unusual characteristics i n the presence of a safety p i l o t who can revert t o normal characteristics. I n such an application, t h e a i r c r a f t equations of motion and s t a b i l i t y characteristics, which a r e estimated by t h e o r e t i c a l calculations and wind-tunnel t e s t i n g , are programed i n t o t h e onboard computers. If t h e desired degree of d e t a i l requires variation i n t h e c h a r a c t e r i s t i c s as a function, say, of airspeed, function generators i n t h e computer make t h i s possible.

CONCLUDING REMARKS A model-controlled simulation technique has been adapted t o a r e l a t i v e l y sophisticated v a r i a b l e - s t a b i l i t y helicopter for study of low-speed handling- q u a l i t i e s requirements. The a b i l i t y of t h e technique t o wash out t h e s t a b i l i t y of the basic helicopter and thus t o command t h e computed response has been demonstrated from analysis of f l i g h t t i m e h i s t o r i e s . Some l a g problems and steady-state e r r o r s were encountered because of l i m i t a t i o n s on t h e maximum error-signal gain which could be achieved. These problems were l a r g e l y over- come, however, by introduction of lead networks which produce t h e correct i n i t i a l response following control inputs and by an i n t e g r a t o r network on t h e rate e r r o r signal which reduces long-term e r r o r s . The r e s u l t s indicate t h a t t h e model technique does, i n f a c t , provide a f e a s i b l e , accurate, and f l e x i b l e approach t o i n - f l i g h t simulation.

REFERENCES 1. Gould, D. G . : The Model-Controlled Method f o r Development of Variable- S t a b i l i t y Aircraft. AGARD Report 402, July 1962.

2. K e l l y , J. R.; and Winston, M. M.: Stability Characteristics of a Tandem- Rotor Transport Helicopter a s Determined by Flight Test. N A S A T N D-2847, 3. Garren, John F.; and Kelly, J. R.: Description of a n Analog Computer Approach t o V/STOL Simulation Employing a Variable S t a b i l i t y Helicopter.

NASA TI9 D-1970, 1964.

Garren, John F. ; Kelly, J. R. ; and Reeder, John P .: A Visual Flight Inves-

4 .

t i g a t i o n of Hovering and Low-Speed VTOL Control Requirements. NASA TN D-2788, 1965.

AGARD Report 408, 5 . Anon. : Recommendations f o r V/STOL Handling Q u a l i t i e s .

October 1962.

6. Garren, John F.; Kelly, J. R . ; and Reeder, John P.: Effects of Gross Changes i n S t a t i c Directional S t a b i l i t y on V/STOL Handling Characteristics NASA TN D-2477, 1964.

Based on a Flight Investigation.

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