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In this paper, the control law design for flexible aircraft is discussed. First, the traditional procedure of decoupling rigid-body and aeroelastic dynamics with low-pass and notch filters is addressed, with focus on controller performance as well as the resulting stability margin issues. A procedure based on a unified formulation of the flexible aircraft dynamics for flight control law design is proposed. In this procedure, the aeroservoelastic dynamics is assessed in the loop, and the offline filtering process is avoided. The formulation is applied to the virtual aircraft generic narrow-body airliner, with improvements in closed-loop performance and stability margins.

Design of next generation aircraft/sensorcraft for improved performance, such as gust load alleviation towards aircraft stability and flutter suppression during its flight operation, may necessitate wing technology that can be controlled and manipulated by active means. Moreover, in recent years the efforts are underway to realize “Fly by Feel” concepts, which are aimed in utilizing on-board sensors (embedded) and actuators (control surfaces) of the aircraft towards the design of active control system for desired performance. This paper presents active control strategies for wings having multiple control surfaces, which are purely based upon in-flight receptance (measured) data. The proposed receptance based control approach has several advantages over the traditional state-space based control because it circumvents approximation errors in reduced order modeling; it captures the true interaction between structure and aerodynamic loads; and it requires modest size of matrices (depending upon the available number of sensors and actuators) for control gain computations. In this study, multi-input state and output feedback control strategies for active aeroelastic control using the method of receptances is developed. The control gains are computed for the extension of flutter boundaries via pole placement. At first, by using numerical receptances obtained from the aeroelastic model of a flexible wing having multiple control surfaces, the proposed methodology is demonstrated. In order to test and demonstrate the receptance method for more complex aircraft geometries, configurations and aerodynamic loading conditions, numerical receptances from Finite Element models of aircraft wings with multiple control surfaces were extracted for the proposed control design. Presented studies and control approach may become the basis for optimal placement and sizing of control surfaces in a given wing section for active aeroelactic control and enhanced flight performance.

A flight control system design technique previously introduced in the literature is applied to four challenging synthesis problems: (1) design of a hypersonic vehicle autopilot, (2) design of a stability augmentation system for a fighter aircraft, including human-in-the-loop desktop simulator evaluation, (3) design of an autopilot for a cruise missile, and (4) design of a control system for an uninhabited rotorcraft. The vehicle models used in the first and third problems have been employed by other researchers using competing design approaches. Human-in-the-loop desktop simulation of the design in the second problem provides a novel evaluation of the resulting stability and command augmentation system’s ability to accommodate system failures from the standpoint of their effect upon operator dynamics. Finally, the ability of the design procedure to accommodate velocity dependent changes in vehicle dynamics is demonstrated in the rotorcraft example The nine-step synthesis procedure to be exercised is based in the frequency domain and evolves from sliding mode control theory.

An approach to the design of damage-tolerant flight control systems is discussed. The design philosophy is based upon a frequency-domain interpretation of pseudo-sliding-mode control. The methodology uses asymptotic observerstoaccommodatetheeffectsofactuatordynamicsinthevehiclebeingcontrolled.Thetechniqueisattractive in that no failure identification or system reconfiguration/adaptation is required after damage. The vehicle model chosen for study is a simple lateral-directional model of an F-18 fighter. The choice of this model was predicated on the ready availability of linearized models of the aircraft at a variety of flight conditions and the fact that an existing robust design was available for performance comparisons. After a brief introduction to the design technique, the F18 design is presented. It is shown that controller scheduling for different flight conditions can be accomplished throughtheobserversalone.Amodelofthehumanpilotisincludedtoassessperformanceandhandlingqualitiesina computer simulation of the resulting pilot/vehicle system. Robustness to modeled damage/failure is demonstrated, including a comparison with an existing robust design applied to the same vehicle model.

Active control of separated flows has become an attractive approach enabling the designer to meet increased performance demands for various systems. Whereas a significant part of the work published so far in the area of closed-loop flow control is based on simulation studies, this paper presents an example of a successful application of a multivariable controller in wind tunnel experiments. A robust H ∞ controller is used to control the spanwise variable reattachment length downstream of a benchmark problem, the backward-facing step. To reduce the conservatism of this approach, a nonlinear static precompensation is included, first. The synthesis of the controller is based on a family of identified linear black-box models, which describe the compensated input/output behavior of the plant. Tracking performance and disturbance rejection of the controller are tested in wind tunnel experiments and shown in this paper.

A mathematical optimization problem is formulated to design several components of an electrical power distribution system for next-generation aircraft. A simple interconnected system consisting of an input e lter and a dc‐dc buck converter is used as the prototype for the optimization demonstration. The components are designed for minimum weight, subject to performance, stability, and peak voltage constraints. Optimized designs for each of the individual subsystems and for the integrated system are presented. It is shown that there is an interaction between thee lterdesign problem and theconverterdesign problem and thatweightimprovementscanbeobtained by considering these interactions. The optimization methodology described enables designers to obtain optimized electric power subsystem designs in a time-efe cient manner.

This paper presents an experimental implementation of an active control system used to suppress flutter in a typical section airfoil. The H2 optimal control system design is based on experimental system identifications of the transfer functions between three measured system variables - pitch, plunge, and flap position - and a single control signal that commands the flap of the airfoil. Closed-loop response of the airfoil demonstrated gust alleviation below the open-loop flutter boundary. In addition, the flutter boundary was extended by 12.4% through the application of active control. Cursory robustness tests demonstrate stable control for variations in flow speed of ± 10%.

Open- and closed-loop tests for a strain-actuated active aeroelastic wing are summarized. Linear quadratic Gaussian (LQG) control laws as well as robust control laws are designed using sensitivity weighted LQG, classical rationalization, and multiple models. Significant vibration suppression and load alleviation are demonstrated, reducing the power spectral density of the first mode's response by an order of magnitude. The flutter dynamic pressure is increased by 12%. The three major performance limitations are the saturation limit of the piezoelectrics, the choice of performance metric or output sensor, and the changes in the dynamic response of the test article. T

The synthesis and experimental validation of a control law for an active flutter suppression system for the active flexible wing wind-tunnel model is presented. The design was accomplished with traditional root locus methods making extensive use of interactive computer graphics tools and simulation-based analysis. The design approach relied on a fundamental understanding of the flutter mechanism to formulate a very simple control law structure resulting in a filter with a "inverted notch" characteristic. This unusual filter characteristic was required to compensate for adverse zero locations in the frequency range near flutter. Wind-tunnel tests of the flutter suppression controller demonstrated simultaneous suppression of two flutter modes, significantly increasing the flutter dynamic pressure. The flutter suppression controller was also successfully operated in combination with a rolling maneuver controller to perform flutter suppression during rapid rolling maneuvers.

This article describes the procedures and results of aeroservothermoelastic studies. The objectives of these studies are to develop the necessary procedures for performing an aeroelastic analysis of an aerodynamically heated vehicle and to analyze a configuration in the classical "cold" state and in a "hot" state. Major tasks include the development of the structural and aerodynamic models, open loop analyses, design of active control laws for improving dynamic responses, and analyses of the closed loop vehicles. The analyses performed focused on flutter speed calculations, short period eigenvalue trends, and statistical analyses of the vehicle response to controls and turbulence. Improving the ride quality of the vehicle, and raising the flutter boundary of the aerodynamically heated vehicle up to that of the cold vehicle, were the objectives of the control law design investigations.

In this paper, Zakian's method of inequalities is applied to the design of a ride-control system for a STOL aircraft. The purpose of the controller is to reduce the normal acceleration experienced by both passengers and crew. The method is based on the synthesis of a controller such that a set of performance specifications and constraints is satisfied. Controllers are designed on the basis of the characteristics of both the closed-loop step response and the closed-loop error response. It is shown that the design of a single-input, single-output controller by the method of inequalities is straightforward, and can be achieved by using a sequence of formulations until the designer is satisfied that no further improvement is necessary.

Through analyses and recent wind-tunnel tests, active flutter suppression has been shown to be a promising technique for preventing wing/external store flutter restrictions. Data measured in the wind tunnel have been used to evaluate the validity of a method for the design and analysis of active feedback control systems. The results of an analytical effort to study the behavior of an active flutter suppression wind-tunnel model are presented and compared with available test data. For this application, the model was aerodynamically represented by subsonic doublet lattice theory and stability was evaluated using modified Nyquist criteria.

This fifth simulation experiment, using the U.S. Navy X-22A variable stability V/STOL aircraft, was undertaken to generate data for the development to flying qualities and advanced flight control system design criteria for the visual shipboard landing task. Since a broad range of flying qualities were simulated, actual landings were not performed because of flight safety considerations. Instead, a discrete three-dimensional position tracking task, presented to the pilot on a Head-Up-Display (HUD), was devised to evoke pilot control and stabilization activity similar to that in the actual shipboard environment. A microwave landing system with precision ranging capability served as the guidance sensor for both the translational rate flight control system mechanizations and for the HUD tracking information. A total of 111 evaluations were performed of various horizontal and vertical translational rate flight control system dynamics. The primary results of the program defined regions of satisfactory and acceptable flying qualities as functions of velocity command gain and the time constant of velocity response. The limits on command gain and time constant for satisfactory flying qualities indicated by this experiment are considerably smaller than those determined in ground simulator experiments. (Author)

It is readily apparent that current longitudinal flying qualities criteria do not adequately account for the effects of dynamic modes introduced by today's complex flight control systems (FCS). To remedy this situation, a combined analytical and experimental investigation was recently conducted using the U. S. Air Force/Cornell Aeronautical Laboratory (USAF/CAL) variable-stability T-33 airplane. Based on an extensive pilot-in-the-loop analysis of the experimental results, a design criterion was developed which is shown to be applicable to a wide range of short-period and FCS dynamics. A simplified version is also presented to provide the designer with preliminary estimates of flying qualities.

To improve ride qualities and structural fatigue life of high-performance aircraft, control of structural flexing is desirable. Linear optimal control via the root-square locus was employed to design a simple, effective bending-control system for four XB-70 coupled longitudinal bending modes. The weighted sum of the mean-square differential angular acceleration and the control surface position was minimized for random wind-gust inputs. Approximation of frequency response characteristics of high-order optimal compensations led to a suboptimal mechanization (fixed-parameter, third-order transfer function with a programmed gain) which produced nearly optimal performance for three diverse low-altitude flight cases. The rms differential angular acceleration was reduced by a factor of 5 for a control expenditure of approximately 13°/sec rms elevon rate per fps of rms wind-gust velocity. Changed vehicle parameters for a fourth flight case require different sensor locations for effective control. Extensive design data show correlation of performance with plant gain and performance-index weighting factors, indicating a potential for predesign prediction of results. Comparable performance is provided by simple, uncompensated feedbacks for some cases, but stability margins are lower than for the near-optimal systems.