Flight investigation of an on-off control for V/STOL aircraft under visual conditions

APPENDIX A SIMULATION TECHNIQUE AND ACCURACY In the signal-flow diagram in figure 12, the simulation techdque and the essential features of the mechanization of the equations of motion f o r the on-off control study are F i g u r e 12.- S i m p l i f i e d signal-flow d i a g r a m for p i t c h axis.

illustrated f o r the longitudinal axis. A corresponding mechanization was used f o r the lateral axis. As indicated in figure 12, motion of the pilot's control stick produced a signal which w a s used to actuate an on-off relay whenever the stick motion exceeded a preselected travel, o r dead band. As indicated in the insert at the top of figure 12, while the stick w a s maintained within the dead band, the relay output w a s zero; hence, was commanded. When the stick was maintained outside the dead band, the no motion relay w a s actuated and its output became plus or minus a constant value, depending on

the direction of the stick displacement. The output of the relay was fed into integrator 0

which computed the model angular velocity on the basis of the control position and the other inputs to the integrator.

The simulation of aerodynamic angular-velocity damping was accomplished by feeding back a signal proportional to the model angular velocity into the summing junction of integrator 0.

Simulation of static stability (i.e., speed stability i n the case of the longitudinal axis) was accomplished by feeding a signal proportional to the forward component of the

aircraft velocity into integrator 0. In the absence of an accurate method for sensing

APPENDIX A airspeed at very low speeds including hovering, ground-referenced velocity signals, which were obtained from an onboard Doppler radar system, were used to provide the required velocity signals.

An electromechanical disturbance generator was used in the tests to provide a con- trolled disturbance level. its character- A description of the disturbance generator and istics is given in appendix B. The disturbance generator permitted the root-mean-square disturbance level to be varied from near z e r o (since the model simulation technique tended t o eliminate external disturbances) to any desired magnitude, T r i m changes of the type which would be caused by a shift in aircraft center-of- gravity position were simulated by feeding a constant voltage into integrator 0.

The degree of simulation accuracy which was obtained during the on-off control investigation can be estimated on the basis of the time histories shown in figure 13. In Acceleration command + - - _I -7

* -..-- oE--- _ _ _ -

a- A _ - ___I - 4 r - Theoretical response Angular velocity, Actual response O r radian/sec , -4 - I I I I I I I ! I . . I . I I I . , I I L L__I 0 4 8 1 2 1 6 20 Time, sec Figure 13.- Comparison of the actual a n d the theoretical responses for MyA/ly = 0.15 radian/sec2.

this figure are shown the acceleration command, the theoretical angular-velocity response, and the actual angular-velocity response for the longitudinal axis. The time histories shown in the figure were obtained for a zero-order model, that is, a model air- craft having neutral stability s o that the response to a control input was pure angular acceleration. The zero-order response was chosen f o r illustration here because it represents the most difficult case for the test helicopter to follow.

By taking the theoretical response in figure 13 as the reference, a time lag of about 0.2 second is observed in the actual response. The amplitude of the actual response, on the other hand, is considered to be reasonably accurate as may be observed by shifting the relative horizontal position of either of the time histories by about 0.2 second.

Despite the fact that time delays on the order of 0.2 second are sometimes considered t o be detrimental t o handling qualities (and perhaps rightfully so), there was no evidence of such problems in these tests. Even though some of the pilots were forewarned of the existence of the time delay, none of them were able t o detect it and it is not believed that this time delay had any adverse effect on the results.

APPENDIX A There are several factors which might account f o r the acceptability of the time delay in the present case. First, the pilot had no cue as to precisely when the control dead band was exceeded since overtravel was available rather than having control stops at the edge of the dead band. Second, the dead band was quite s m a l l so that full control was commanded in a v e r y small control travel and, hence, in a shorter time than the pilot normally takes to put in a large control deflection with a proportional control.

Therefore, the total elapsed time f r o m when the pilot senses the need f o r a control dis- placement to the time when the aircraft has a perceivable response is probably consid- erably shorter in the present simulation than in aircraft with conventional control sys- tems. Also, in any aircraft there is some time delay in the response t o control deflec- tion. Third, a factor which might mask the effects of the time delay in the present case is that, on a long-term basis (greater than a few tenths of a second), the attitude e r r o r signal s e e s an e r r o r due to the lag and feeds in additional control so that the attitude change after a sufficient time will be nearly correct. It is seen f r o m this last example that the time delay associated with this simulation lacks a clear analogy with classical time delays in that, on a long-term basis, the simulation technique used tends to over- control to compensate f o r the initial delay. All things considered, the results presented herein would perhaps correspond most closely to an aircraft with a control system having a combined first-order and transport time delay between 0.1 and 0.15 second.

APPENDIX B DESCRIPTION OF DISTURBANCE GENERATOR The disturbance generator used in the present investigation is an electromechanical device which employs the voltage outputs of three cam-driven potentiometers t o provide random inputs t o the model aircraft. Each of the three cams rotates at a slightly differ- ent frequency near 1 cycle p e r minute. The voltage output of each of the potentiometers is combined into a single output so as to obtain an output which repeats infrequently (owing to the slight variation in cam speeds). By superimposing the outputs of the cam- driven potentiometers on different fixed gains, it was possible to achieve somewhat inde- pendent sources of disturbances f o r the different aircraft axes.

A sample of the disturbance generator output is shown in figure 14. The maximum voltage peaks correspond to about 2 1 to 3 times the root-mean-square output of the gen- erator which is in reasonable agreement with the familiar rule of thumb for natural tur- bulence. A power-spectral analysis of the simulated disturbances is shown in figure 1 5 for root-mean-square disturbance level of 0.04 rad/sec2 for which case the peak disturb- ances were about 0.1 rad/sec2. As can be observed from this plot, the intensity of the disturbances was concentrated at f a i r l y low frequencies which would necessitate positive corrective action on the part of the pilot.

r

20 40 60 80 100 120 Time, sec Figure 14.- Sample of disturbance-generator output. Root-mean-square level = -lV.

2 1

APPENDIX B

.OIOC .OOOl 10.0 1 . 0 0. I Frequency, radion/sec Figure 15.- Power spectral analysis of simulated disturbances scaled to provide a selected root-mean-square disturbance.

2 2

APPENDIX C DESCRIPTION OF TASKS Maneuvering Tasks Longitudinal quick starts and stops.- During the longitudinal quick starts and stops, - ~ - _ _ _ _ _ _ _ _ - - - the helicopter was pitched down to accelerate to about 4 5 knots. After a short run at 4 5 knots, the aircraft was rapidly decelerated to a hover over a preselected spot.

Lateral quick starts and stops.- During the lateral quick starts and stops, the helicopter was rolled t o the right o r left to obtain a lateral acceleration until a sideward velocity of about 25 knots w a s attained. The aircraft was then rolled in the opposite direction and rapidly decelerated t o a hover over a preselected point.

Turn reversals and landing approach.- The pilot made visual approaches to a spe- cific landing spot at a speed of about 4 5 knots and an altitude of approximately 100 feet (30.5 m). During this approach the pilot executed rapid S-turn maneuvers, both to the left and to the right, before coming to a hover and descending t o the intended landing point.

Precision Tasks

Precision hovering. - The precision hovering task involved attempts at proceeding a

very short distance to a point over a spot on the ground, stopping, and accurately main- taining this position. In order to provide a close visual reference, the task was per- formed a t heights no greater than 20 feet (6.1 m).

Vertical landing. - After gaining reasonable familiarity with the test conditions during the precision hovering task, vertical landings to touchdown were attempted.

Precision ground track. - The precision ground-track task involved attempts at

This task was performed at a height flying in a straight line at an airspeed of 4 5 knots.

of approximately 100 feet (30.5 m).

REFERENCES 1. Lollar, Thomas E.; Bus, Frank J.; and Dolliver, David M.: Control Requirements and Control Methods f o r Large V/STOL Aircraft. p r e p r i n q 650808, SOC.Automotive Engrs., Oct. 1965.

2. Garren, John F., Jr.; and Kelly, James R. : Description of an Analog Computer a Variable-Stability Helicopter.

Approach t o V/STOL Simulation Employing NASA TN D-1970, 1964.

3. Garren, John F., Jr.; and Kelly, J a m e s R.: Application of the Model Simulation Technique to a Variable-Stability Helicopter for Simulation of VTOL Handling Presented at the 27th Meeting of the AGARD Flight Mechanics Panel Qualities.

(Rome, Italy), Oct. 11-12, 1965.

4. Garren, John F., Jr.; Kelly, J a m e s R.; and Reeder, John P.: A Visual Flight Investi- gation of Hovering and Low-Speed Control Requirements. NASA T N D-2788, 1965.

5. Besco, Robert 0. : Handling Qualities Criteria for Manned Spacecraft Attitude Control Systems. AIAA/NASA Third Manned Space Flight Meeting, CP-10, Am. Inst.

Aeron. Astronaut., Nov. 1964, pp. 271-275.

NASA-Langley, 1966 L-4985

I

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