Pilot-induced Oscillations Research Articles

Simulator Platform Motion Effects on Pilot-Induced Oscillation Prediction

Simulator motion platform characteristics were examined to determine if the amount of motion affects pilotinduced oscillation prediction. Five test pilots evaluated how susceptible 18 different setsof pitch dynamics were to pilot-induced oscillations with three different levels of simulation motion platform displacement: large, small, and none. The pitch dynamics were those of a previous in-e ight experiment, some of which elicited oscillations. These in-e ight resultsserved as truth data for the simulation. As such, the in-e ight experiment wasreplicated asmuch as possible. Objective and subjective data were collected and analyzed. With large motion, pilot-induced oscillation and handling qualities ratings matched the e ight data more closely than with small motion or no motion. Also, regardlessof the aircraft dynamics, large motion increased pilot cone dencein assigning handling qualitiesratings, reduced safety pilot trips, and lowered touchdown velocities. Whereas both large and small motion provided a pitch rate cue of high e delity, only large motion presented the pilot with a high e delity vertical acceleration cue.

Robust flight control design with handling qualities constraints using scheduled linear dynamic inversion and loop-shaping

A technique for obtaining a full-envelope decoupled linear flight control design is presented. The methodology begins with a reduced-order linear dynamic-inversion technique that is scheduled over the flight envelope. The reduced order dynamic inverter can offer a significant reduction in the number of state variables to be sensed or estimated as compared to typical applications of inverse dynamic control. The technique can provide desired input-output characteristics including control decoupling. The required gain scheduling of the reduced order dynamic inversion is straightforward. Uncertainty is introduced by perturbing the stability derivatives in the vehicle model at each of the flight conditions considered. The effects of uncertainty are then reduced by additional feedback loops involving a diagonal compensation matrix obtained through application of a loop shaping procedure based upon a quantitative feedback theory predesign technique. The tendency of quantitative feedback theory to produce high-bandwidth conservative designs is mitigated by the scheduling and decoupling associated with the dynamic inversion. Finally, handling qualities and pilot-induced oscillation tendencies are evaluated using a structural model of the human pilot implemented in an interactive computer program that can include the effects of nuisance nonlinearities such as actuator saturation. The proposed methodology is applied to the design of a lateral-directional flight control system for a piloted supermaneuvarable fighter aircraft.

Aircraft Handling Qualities and Pilot-Induced Oscillation Tendencies with Actuator Saturation

A pilot/vehicle analysis technique is offered that can provide predictions of handling-qualities levels and pilotinduced oscillation rating levels in single-axis tracking tasks, with linear or nonlinear vehicle dynamics. The technique is applied to the three vehicle conŽ gurations evaluated in the series of  ight tests sponsored by the U.S. Air Force that are identiŽ ed as HAVE LIMITS. The ability of the analysis technique to provide excellent quantitative predictions of the effects of actuator rate saturation on the handling qualities and pilot-induced oscillation susceptibility of the aircraft conŽ gurations is demonstrated. An interactive computer code is discussed that automates much of the pilot/vehicle analysis.

Overcoming pilot-induced oscillations in the space shuttle

Overcoming pilot-induced oscillations in the space shuttle

Development of the Probe-and-Drogue Handling Qualities Demonstration Maneuver

The development of the probe-and-drogueaerial refueling task as a handlingqualities demonstrationmaneuver is described. The probe-and-drogue task is an excellent test of handling qualities for precision  ying and is also a superb check for pilot-induced oscillation tendencies. The greatest difŽ culty with this maneuver has been in the deŽ nitionof the task itself including the deŽ nitions of speciŽ c performance limits. The task described herein differs from that described in the Notice of Change to Flying Qualities of Piloted Aircraft (MIL-STD-1797A) in that it requires precise drogue tracking in addition to a percentage of successful hookups. A task representative of that deŽ ned in MIL-STD-1797A that required at least three hookups in six attempts for desired performance was used as the evaluationmethod for theHAVEGAS project conducted by the U.S.Air Force Test Pilot School. This version of the maneuver, however, was found to be too easy even for poorly rated airplanes. Revised maneuver descriptions and performance requirements were developed and evaluated in the HAVE GAS II project that was undertaken as part of the U.S. Air Force Demonstration Maneuvers program and again conducted by the U.S. Air Force Test Pilot School. Three candidate tasks were evaluated in this test program: 1) the probe-and-drogue hookup task, 2) a new drogue tracking task, and 3) a new drogue pointing task. Although the drogue tracking task was found to be best for evaluating closed-loop handling qualities, it was limited by the fact that it would not reveal potential problems such as bow-wave effects that may be encountered within 5 ft of the basket. Thus it was concluded that a combination of the drogue tracking task and the probe-and-drogue hookup task should be  own to thoroughly evaluate closed-loop handling qualities in the aerial refueling mission.

Assessing Aircraft Susceptibility to Nonlinear Aircraft-Pilot Coupling/Pilot-Induced Oscillations

A unified approach for assessing aircraft susceptibility to aircraft-pilot coupling (or pilot-induced oscillations) which was previously reported in the literature and applied to linear systems is extended to nonlinear systems, with emphasis upon vehicles with actuator rate saturation. The linear methodology provided a tool for predicting: (1) handling qualities levels, (2) pilot-induced oscillation rating levels and (3) a frequency range in which pilot-induced oscillations are likely to occur. The extension to nonlinear systems provides a methodology for predicting the latter two quantities. Eight examples are presented to illustrate the use of the technique. The dearth of experimental flight-test data involving systematic variation and assessment of the effects of actuator rate limits presently prevents a more thorough evaluation of the methodology.

Pilot-Induced Oscillations Involving Multiple Nonlinearities

A pilot-induced oscillation is a type of instability caused by dynamic couplingbetween the pilot and the aircraft. Linear system analysis reveals that the pilot/aircraft system is unstable during the oscillation event. However, nonlinearities in the feedback system can limit the response variables so that a sustained, constant amplitude limit cycle results. A method is presented to analyze the limit cycle behavior of pilot/aircraft systems as a means of understanding pilot-inducedoscillations and their causes. A new computationaltechnique is considered and a new methodof testing limit cycle stability is presented. An analysisof the YF-12 pilot-inducedoscillation characteristics is used to demonstrate the method.

Robust, Decoupled, Flight Control Design with Rate-Saturating Actuators

Techniques for the design of control systems for manually controlled, high-performance aircraft must provide the following: (1) multi-input, multi-output (MIMO) solutions, (2) acceptable handling qualities including no tendencies for pilot-induced oscillations, (3) a tractable approach for compensator design, (4) performance and stability robustness in the presence of significant plant uncertainty, and (5) performance and stability robustness in the presence actuator saturation (particularly rate saturation). A design technique built upon Quantitative Feedback Theory is offered as a candidate methodology which can provide flight control systems meeting these requirements, and do so over a considerable part of the flight envelope. An example utilizing a simplified model of a supermaneuverable fighter aircraft demonstrates the proposed design methodology.

Theory for Roll-Ratchet Phenomenon in High-Performance Aircraft

Roll-ratchet refers to a high frequency oscillation which can occur in pilot-in-the-loop control of roll attitude in high performance aircraft. The frequencies of oscillation are typically well beyond those associated with the more familiar pilot-induced oscillation. A structural model of the human pilot which has been employed to provide a unified theory for aircraft handling qualities and pilot-induced oscillations is employed here to provide a theory for the existence of roll-ratchet. It is hypothesized and demonstrated using the structural model that the pilot's inappropriate use of vestibular acceleration feedback can cause this phenomenon, a possibility which has been discussed previously by other researchers. The possible influence of biodynamic feedback on roll ratchet is also discussed.

Unified Theory for Aircraft Handling Qualities and Adverse Aircraft-Pilot Coupling

A unified theory for aircraft handling qualities and adverse aircraft-pilot coupling or pilot-induced oscillations is introduced. The theory is based on a structural model of the human pilot. A methodology is presented for the prediction of (1) handling qualities levels; (2) pilot-induced oscillation rating levels; and (3) a frequency range in which pilot-induced oscillations are likely to occur. Although the dynamics of the force-feel system of the cockpit inceptor is included, the methodology will not account for effects attributable to control sensitivity and is limited to single-axis tasks and, at present, to linear vehicle models. The theory is derived from the feedback topology of the structural model and an examination of flight test results for 32 aircraft configurations simulated by the U.S. Air Force/CALSPAN NT-33A and Total In-Flight Simulator variable stability aircraft. An extension to nonlinear vehicle dynamics such as that encountered with actuator saturation is discussed.

Nonlinear Longitudinal Stability of a Symmetric Aircraft

The longitudinal stability of a symmetric aircraft is examined beyond the well-known linearization about a steady mean state, leading to the phugoid and short-period modes. The exact equations of longitudinal motion of a symmetric aircraft are considered, i.e., balance of longitudinal and transverse force (without side force ) and balance of pitching moment; the drag terms included are friction and lift-induced drag, plus nonsymmetric lift ‐ drag polar; the mass density is taken as a constant, as well as thrust along the e ight path. Elimination would lead to a fourth-order nonlinear differential equation for the angle of attack relative to the angle of zero pitching moment, if the acceleration of the e ight-path angle is neglected, and the e ight-path angle is moderate. In the case of a small e ight-path angle, it simplie es to a thirdorder differential equation, containing a set of nonlinear corrections, to a second-order linear equation, specifying sinusoidal oscillation of the relative angle of attack. It is shown, by a small perturbation method, that forced oscillations occur at its harmonics (viz. double or triple frequency ), and free oscillations can have decaying or growing amplitude. In the case of statically stable aircraft, the oscillations have a short period and grow or decay slowly. In the case of statically unstable aircraft the growth is rapid, as in a pilot-induced oscillation.

Flight Control System Design with Rate Saturating Actuators

Actuator rate saturation is an important factor adversely affecting the stability and performance of aircraft flight control systems. It has been identified as a catalyst in pilot-induced oscillations, some of which have been catastrophic. A simple design technique is described that utilizes software rate limiters to improve the performance of control systems operating in the presence of actuator rate saturation. As described, the technique requires control effectors to be ganged such that any effector is driven by only a single compensated error signal. Using an analysis of the steady-state behavior of the system, requirements are placed upon the type of the loop transmissions and compensators in the proposed technique. Application of the technique to the design of a multi-input/multi-output, lateral-directional control system for a simple model of a high-performance fighter is demonstrated as are the stability and performance improvements that can accrue with the technique.

Pilot-Induced Oscillation Analysis and Prediction with Actuator Rate Limiting

Akeyfeatureofall of therecentwell-recordedseverepilot-induced oscillations (PIOs)isthepresenceof actuator rate limiting. As part of attempts to develop a comprehensive understanding of PIO phenomena, the results of an effort focused on improved understanding and analysis of actuator rate limiting are presented. For this effort describing functions were used in concert with modern computer simulation techniques to quantify the added phase lag, magnitude reduction, and bandwidth reduction of a rate-limited actuator in terms of the input and actuator design parameters. The results and inverse describing function techniques are then employed to analyze thelimit cyclepotential of a system that isessentially linearexceptfortheseriesnonlinearity. Thewell-documented X-15 landing e are PIO is used to exemplify these analysis techniques. From this example it is shown that these techniques can be used to provide a prediction of the limit cycle or PIO frequency.

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Pilot-Induced Oscillation (PIO) was encountered in the company flight test of the T-2CCV research vehicle. The main cause of the roll PIO was due to unexpected large pilot input and to increased delay of the aircraft response by its large pilot input. The delay of roll response has much difference between with large input and with small input. That is why the flight control system falls into nonlinear system by actuator rate limit in case of controlling the aircraft with large pilot input. This paper shows analysis of the characteristics of the PIO and its improvement.

Synthesis and evaluation of an H2 control law for a hovering helicopter

The design and simulator evaluation of a rate-command flight-control law for a UH-60 helicopter in near-hover flight conditions are described. The multivariable control law was synthesized using an //2 method, which, through weighting functions, directly shapes the singular values of the sensitivity (/ + GK)~l, comple- mentary sensitivity GK(I + GK)~ l, and control input K(I + GK)~l transfer-function matrices. The design was implemented on the vertical motion simulator, and four low-speed hover tasks were used to evaluate the control system characteristics. The pilot comments indicated good decoupling and quick response characteristics, but also revealed a mild pilot-induced oscillation tendency in the roll axis. HE purpose of the research documented in this paper is to investigate the application of an H2 multivariable de- sign method to synthesize flight-control laws for helicopters. The objective of the work is to determine through analysis and piloted simulation the strengths and weaknesses of the H2 method within the context of a modern rotorcraft design specification. The choice of a multivariable design method is natural to the problem presented because of the highly coupled nature of helicopter dynamics and the desired level of feedback dictated by modern design specifications. Helicopter flight-control laws designed through classical single-input, single-output techniques are satisfactory when requirements do not infringe on the frequency range of significant interaxis cross coupling due to the rotor. Previous generation helicopters and their design requirements have complied with this assumption. However, future helicopters will use hingeless and bearingless rotors with more complex coupling, and these helicopters will have control-law design requirements demanding higher levels of feedback. To ensure that the design criteria are sufficiently stringent to represent future requirements, the control law here was de- signed according to the Section 3 hover requirements in the Aeronautical Design Standard (ADS-33C).1 Previous work on multivariable techniques for rotorcraft flight-control design2'4 complied with part or none of the ADS-33C requirements, making these works difficult to interpret in terms of this new standard. These previous works also used a low-order model (usually a 6-degree-of-freedom model) to design the flight-control law, and the final design was usually checked on a higher-order model in analysis or simulation. These higher-order models did not fully represent the rotor-fuselage dynamics, lacking either the coupled nature of the rotor dynamics (between the fuselage and the rotor and/or between the rotor degrees of freedom) or lacking the rotor lead-lag degree of freedom. These effects are known limiting factors in the feedback gains of helicopter flight-control laws and are represented in the analysis and piloted simulator model here. The importance of these effects was illustrated in advanced digital optical control system (ADOCS) research, where feedback gain reductions of 40% were necessary to avoid lead-lag instabilities.5 This paper continues with a description of the models used in the design process and piloted simulation, followed by a description of the method used to synthesize the flight-control law. The salient features of a rate-command control-law de- sign are then discussed, followed by descriptions of the simu- lation environment and task arrangement. Finally, the results of the simulation are discussed and conclusions drawn. More details on the design method, design models, and flight-con- trol design can be found in Ref. 6, whereas more details related to the simulation experiment can be located in Ref. 7.

Roll metrics for handling qualities - The landing approach

Introduction H IGHLY augmented aircraft have shown handling qualities degradation in the roll axis due to increased flight control system augmentation and high-order effects. In a recent paper, alternate metrics were proposed, based on Gibson's method, to identify handling qualities deficiencies, and level 1 boundaries were established for the tracking task. In the present paper, new guidelines for the landing task are presented and validated, thus increasing the confidence level of the methodology as a guideline for designing good handling qualities as well as for their assessment. The landing task is normally identified as compensatory but, unlike tracking, is essentially a low-gain task with the pilot's control activity centered around a fixed-phase lag value of about 120 deg (Ref. 3). The analysis of the landing task is based on the LATHOS experiment data base relative to class IV, category C aircraft and involves both instrument landing as well as visual landing tasks. In all, 16 configurations were analyzed with different roll time constant rr, pure time delay Tdt and roll control effectiveness. A detailed description of the configurations can be found in Ref. 4. The roll metrics defined for this analysis are the effective time delay Tc{f (system delay in response to a change in forcing function as seen by the pilot), roll control effectiveness LFas pss/rr, and acceleration ratio Ra defined as the ratio between maximum and initial roll accelerations and given by Ra = max/(LFas). The next two sections present the results of the analysis, level 1 boundaries are identified as well as pilot-induced oscillations (PIO) and ratcheting-prone configurations using the roll metrics just described.

Roll-performance criteria for high augmented aircraft

Highly augmented fly-by-wire aircraft have shown a degradation in roll performance during compensatory tasks such as tracking, similar to the ones observed for the pitch axis. Gibson's method, which involves a combination of time-domain and frequency-domain techniques, has been shown to provide excellent handlingqualities criteria for the pitch axis, especially in identifying pilot-induced oscillations. The present work extends Gibson's method to develop similar handling-qualities criteria for the roll-axis control system. The effects of parameters such as roll damping, transport delay, system gain, and corresponding pilot ratings are studied for developing handling-qualities criteria and to evaluate the flight-control system performance. Flying-qualities levels are obtained in both the time and the frequency domain. In addition, frequency-domain analysis is used to identify pilot-induced oscillation-prone configurations and the presence of roll ratcheting. The analysis is performed and validated using an existing experimental data base on lateral flying qualities of highly augmented class IV aircraft.

ANALYSIS OF PILOT INDUCED OSCILLATIONS

Ever since manually controlled flight of powered aircrafts, pilots have experienced pilot induced oscillations (PIO). PIO problem is one in which an oscillation of the aircraft occurs, that is difficult or even impossible for the pilot to stop. A central characteristic of the phenomenon is that the airplane is stable both stick-fixed or stick-free, hence the name pilot induced oscillation. The presented work studies the effect of pilot and aircraft parameters on pitch tracking performance. Also conditions that bring the pilot-aircraft system to instability were discussed. An appropriate pilot model convenient for PIO analysis is selected and a closed loop analysis of pilot-control system-aircraft has been done in time domain. The pilot dynamics was found to be the most significant factor of PIO. When the pilot model includes neuromuscular system dynamics, the pilot gain required to sustain an oscillation was found to be. 55% of that value with the pilot as a pure gain.

Technique for predicting longitudinal pilot-induced oscillations

A technique for predicting the susceptibility of an aircraft to longitudinal pilot-induced oscillations (PIOs) is proposed. The technique employs the optimal control model (OCM) of the human pilot formulated for pitch attitude command tracking tasks. The criterion is very simple and is based solely on the characteristics of the pilot/vehicle open-loop transfer function as predicted by the OCM. The criterion has its genesis in the results of a previous study using the OCM in pitch attitude tracking tasks and was evaluated here using results from flight tests of 25 configurations flown on the NT-33 variable stability aircraft and three additional configurations from flight tests of a high-performance aircraft. Only the index of performance weighting coefficients in the OCM are considered as model variables, and a technique for generating these coefficients is reviewed.

Loop separation parameter - A new metric for landing flying qualities

Loop Separation Parameter is a metric used to evaluate the longitudinal flying qualities of aircraft in the landing flight phase. It combines simple pilot models with classical root locus and frequency response methods to predict Cooper-Harper pilot ratings for aircraft in the landing phase. Loop Separation Parameter was originally developed using fighter-type aircraft flight test data; in this paper the concept is extended and shown to be a good tool for predicting pilot ratings for large bomber, transport, and highly augmented Space Shuttle configurations. Loop Separation Parameter anticipates the tendency for and predicts the frequency of pilot-induced oscillations. The effects of control system time delay and altitude lag time constant are discussed. It is recommended that Loop Separation Parameter be included in the Flying Qualities Handbook. .