Attitude Control Of Flexible Spacecraft Research Articles

Rebuttal to “Fault-tolerant sliding mode attitude control for flexible spacecraft under loss of actuator effectiveness”

A paper, “Fault-tolerant sliding mode attitude control for flexible spacecraft under loss of actuator effectiveness,” was published in the journal Nonlinear Dynamics. The Lyapunov proof is flawed and does not prove stability for the system. It is also unclear how the simulation results in the paper were obtained using the control law. This rebuttal will demonstrate the flaws in the original paper.

Authors’ response to the rebuttal by A. Cushman

This is a response to the rebuttal by A. Cushman regarding our paper “Fault-tolerant sliding mode attitude control for flexible spacecraft under loss of actuator effectiveness” (Nonlinear Dyn. doi:10.1007/s11071-011-0007-5, 2011).

Simultaneous Vibration Reduction and Attitude Control of Flexible Spacecraft with On-Off Actuators

Abstract This paper deals with the problem of simultaneous vibration suppression and robust attitude control of orbiting spacecraft with flexible appendages. Based on integral vaziable structure control theory, a discontinuous attitude control law is fustly derived to achieve the desired position of the spacecraft, taking explicitly into account external disturbance and nonlinearity. To reconstruct estimates of the system states for use in a full information variable structure control law, an asymptotic variable structure observer is also employed and the fulfillment of sliding condition, including the case when estimated states aze used, is also verified. To achieve the residual vibration free, the shaped input attenuator is also employed after the end of input in an open-loop manner. In addition, the system also employs the pulse-width modulation technique to achieve the desired control torque signals. The performance and efficacy of the proposed control scheme are illustrated with a compazison of diffe...

A Robust Nonlinear Control Approach for Tip Position Tracking of Flexible Spacecraft

In this paper, the problem of attitude control of a 3D nonlinear flexible spacecraft is investigated. Two nonlinear controllers are presented. The first controller is based on dynamic inversion, while the second approach is composed of dynamic inversion and γ-synthesis schemes. The extension of dynamic inversion approach to flexible spacecraft is impeded by the nonminimum phase characteristics when the panel tip position is taken as the output of the system. To overcome this problem, the controllers are designed by utilizing the modified output redefinition approach. It is assumed that only three torques in three directions on the hub are used. Actuator saturation is also considered in the design of controllers. To evaluate the performance of the proposed controllers, an extensive number of simulations on a nonlinear model of the spacecraft are performed. The performances of the proposed controllers are compared in terms of nominal performance, robustness to uncertainties, vibration suppression of panel, sensitivity to measurement noise, environment disturbance, and nonlinearity in large maneuvers. Simulation results confirm the ability of the proposed controller in tracking the attitude trajectory while damping the panel vibration. It is also verified that the perturbations, environment disturbances, and measurement errors have only slight effects on the tracking and damping performances.

Robust attitude and vibration control of a nonlinear flexible spacecraft

AbstractIn this paper, the problem of attitude control of a three dimension nonlinear flexible spacecraft is investigated. Two nonlinear controllers are presented. The first controller is based on dynamic inversion, while the second approach is composed of dynamic inversion and µ‐synthesis schemes. It is assumed that only three torques in three directions on the hub are used. Actuator saturation is also considered in the design of controllers. To evaluate the performance of the proposed controllers, an extensive number of simulations on a nonlinear model of the spacecraft are performed. The performances of the proposed controllers are compared in terms of nominal performance, robustness to uncertainties, vibration suppression of panel, sensitivity to measurement noise, environmental disturbance and nonlinearity in large maneuvers. Simulation results confirm the ability of the proposed controller in tracking the attitude trajectory while suppressing the panel vibration. It is also verified that the perturbations, environment disturbances and measurement errors have only slight effects on the tracking and suppression performances.Copyright © 2010 John Wiley and Sons Asia Pte Ltd and Chinese Automatic Control Society

Fuzzy Adaptive Variable Structure Active Attitude Control of Flexible Spacecraft

This paper presents a novel control system design method for the three-axis-rotational tracking and vibration stabilization of a spacecraft with flexible appendages. Based on the sliding control theory, a robust attitude control law is derived to control the attitude motion of spacecraft. For actively suppressing the induced vibration, both fuzzy methods and strain rate feedback (SRF) control methods are presented. Numerical simulations are performed to show the feasibility and effeteness of the proposed methods.

Fault-tolerant sliding mode attitude control for flexible spacecraft under loss of actuator effectiveness

In this paper, a novel fault-tolerant attitude control synthesis is carried out for a flexible spacecraft subject to actuator faults and uncertain inertia parameters. Based on the sliding mode control, a fault-tolerant control law for the attitude stabilization is first derived to protect against the partial loss of actuator effectiveness. Then the result is extended to address the problem that the actual output of the actuators is constrained. It is shown that the presented controller can accommodate the actuator faults, even while rejecting external disturbances. Moreover, the developed control law can rigorously enforce actuator-magnitude constraints. An additional advantage of the proposed fault-tolerant control strategy is that the control design does not require a fault detection and isolation mechanism to detect, separate, and identify the actuator faults on-line; the knowledge of certain bounds on the effectiveness factors of the actuator is not used via the adaptive estimate method. The associated stability proof is constructive and accomplished by the development of the Lyapunov function candidate, which shows that the attitude orientation and angular velocity will globally asymptotically converge to zero. Numerical simulation results are also presented which not only highlight the ensured closed-loop performance benefits from the control law derived here, but also illustrate its superior fault tolerance and robustness in the face of external disturbances when compared with the conventional approaches for spacecraft attitude stabilization control.

Robust Control of Flexible Mechanical Systems by Utilizing Symmetry and Its Application to Large Space Structures

The control of flexible mechanical systems has been extensively studied in a variety of fields. For mechanical systems with collocated sensors and actuators, the use of symmetric controllers is known to be effective. The symmetric controller guarantees the robust stability of the closed-loop system irrespective of the system parameters by virtue of the structure of the dynamical equation having positive definite or semidefinite symmetric coefficient matrices. Although this controller is superior to existing parametric approaches, systematic procedures for obtaining specific control performance other than the stability of the closed-loop system have not yet been fully established. This paper proposes an optimal design method for symmetric controllers in the sense of the H infin norm by solving the linear matrix inequalities reduced from the bounded real lemma under the symmetry constraint. This method is then extended to static output feedback control and gain-scheduled control. Finally, the design method is applied to a flexible spacecraft attitude control problem, and its validity is confirmed.

An Attitude Control of Flexible Spacecraft Using Fuzzy-PID Controller

This primary objective of this study is to demonstrate simulation and ground-based experiment for the attitude control of flexible spacecraft. A typical spacecraft structure consists of the rigid body and flexible appendages which are large flexible solar panels, parabolic antennas built from light materials in order to reduce their weight. Therefore the attitude control has a big problem because these appendages induce structural vibration under the excitation of external forces. A single-axis rotational simulator with a flexible arm is constructed with on-off air thrusters and reaction wheel as actuation. The simulator is also equipped with payload pointing capability by simultaneous thruster and DC servo motor actuation. The experiment of flexible spacecraft attitude control is performed using only the reaction wheel. Using the reaction wheel the performance of the fuzzy-PID controller is illustrated by simulation and experimental results for a single-axis rotational simulator.

Attitude manoeuvring and vibration reducing control of flexible spacecraft using smart materials

The question of robust attitude control and vibration reduction of flexible spacecraft during manoeuvre is considered. The design approach presented here treats the problem of spacecraft attitude control separately from the elastic vibration suppression, by designing a dynamic controller, using attitude measures alone for desired attitude rotation and smart materials for active vibration suppression. The proposed attitude controller does not need the knowledge of modal variables and spacecraft angular velocity, which is compensated by appropriate dynamics of the controller. For actively suppressing the vibrations introduced to the flexible appendages, embedded piezoelectric ceramic patches are used as both sensors and actuators, to detect and counter-react the induced vibration during and after the manoeuvre. Here, a modified positive position feedback (PPF) control strategy, targeting at the first three flexible modes of the flexible spacecraft is taken into account. The performances of this approach are assessed in terms of the level of vibration reduction at resonance modes and the speed of response. For comparison, the proportional plus derivative control strategy is adopted and compared with proposed control approach. Simulation results demonstrate the effectiveness of the controller design strategy for attitude control and the modified PPF compensators for active vibration suppression.

Adaptive Sliding Mode Attitude and Vibration Control of Flexible Spacecraft Under Unknown Disturbance and Uncertainty

This paper is concerned with the development of a control system for rotational maneuver and vibration suppression of a flexible spacecraft. It is assumed that the system parameters are unknown. The design approach presented here treats the problem of spacecraft attitude control separately from the elastic vibration suppression problem. As a stepping stone, a state feedback sliding mode control command is designed for the reaction wheel to achieve the reference trajectory tracking control of attitude angle. This is followed by the design of an adaptive sliding mode control (ASMC) law using only the output for robust stabilization of spacecraft in the presence of parametric uncertainty and external disturbances. Although this controller has the ability to reject the disturbance, deal with uncertainty and to ensure that the system output errors asymptotically converge to the sliding mode during the commanded motion, it excites the elastic modes of flexible appendages. The undesirable vibration is actively suppressed by applying feedback control voltages to the piezoceramic actuators, in which the control voltages are determined using the modal velocity feedback control method. Both analytical and numerical results are presented to show the theoretical and practical merits of this hybrid approach

Optimal H infinity PD Controller Design for Flexible Spacecraft by Branch and Bound Algorithm

This paper proposes optimization of the PD controller gains by minimizing H infinity norm for flexible spacecraft attitude control. Solving the problem is known to be difficult because it is described by a bilinear matrix inequality (BMI) that cannot be transformed into linear matrix inequalities (LMIs). In order to solve it within practical computation time, we revise the branch and bound algorithm and incorporate the quantization of the variable domains. The ETS-VI mathematical model is used to demonstrate its ability.

クォータニオンフィードバックによる柔軟宇宙機の姿勢制御

クォータニオンフィードバックによる柔軟宇宙機の姿勢制御

Three-axis Attitude Control for Flexible Spacecraft by Lyapunov Approach under Gravity Potential

Attitude control law synthesis for the three-axis attitude maneuver of a flexible spacecraft model is presented in this study. The basic idea is motivated by previous works for the extension into a more general case. The new case includes gravitational gradient torque which has significant effect on a wide range of low earth orbit missions. As the first step, the fully nonlinear dynamic equations of motion are derived including gravitational gradient. The control law design based upon the Lyapunov approach is attempted. The Lyapunov function consists of a weighted combination of system kinetic and potential energy. Then, a set of stabilizing control law is derived from the basic Lyapunov stability theory. The new control law is therefore in a general form partially validating the previous work in some sense.

Passive Attitude Control of Flexible Spacecraft from Quaternion Measurements

In this work, we propose a dynamic controller for a spacecraft with flexible appendages and based on attitude measurements. This control ensures the asymptotic fulfillment of the objectives in the case of rest-to-rest maneuvers when a failure occurs on the accelerometer sensors, so that the angular velocity is not available for feedback. Also, it is assumed that the modal variables describing the flexible elements are not measured. This is a lower level controller and is to be selected at the higher level by a supervisor when an emergency situation is detected.

A Suggestion of Fractional-Order Controller for Flexible Spacecraft Attitude Control

A controller design method for flexible spacecraft attitude control isproposed. The system is first described by partial differential equationwith internal damping. Then the frequency response is analyzed, and thethree basic characteristics of the flexible system, namely, averagefunction, lower bound and upper bound are defined. On this basis, afractional-order controller is proposed, which functions as phasestabilization control for lower frequency and smoothly enters toamplitude stabilization at higher frequency by proper amplitudeattenuation. It is shown that the equivalent damping ratio increases inproportion to the square of frequency.

Nonlinear Synthesis, Optimization and Design of the Spacecraft Gyromoment Attitude Control Systems

The optimization methods for a high-accuracy nonlinear feedback attitude control of the large-scale flexible spacecraft under their spatial rotation manoeuvers are described. The effective damping of oscillations within such flexible structures, using gyrodynes (single gimbal control moment gyroscopes) with electromagnetic suspended rotors, are developed. Some mathematical methods, software for computer-aided design and results are presented.

Multiaxis tracking and attitude control of flexible spacecraft with reaction jets

The objective of this paper is to show that integral pulse frequency modulation (IPFM) reaction jet controllers can perform, with precision: multiaxis tracking of moving objects, attitude control of flexible spacecraft under orbit-adjustment forces, single-axis slews, and vibration suppression. Two sets of position, rate, and acceleration command profiles for multiaxis tracking are presented: one for a payload initially facing the zenith and the commands based on a pitch-roll sequence; the other for a payload facing the nadir and the commands based on a roll-pitch sequence. The procedure for designing an IPFM controller for tracking a given, inertial acceleration command profile is summarized. Important elastic modes of a spacecraft are identified according to their spontaneous attitude and rate response to the minimum impulse bit of the thrusters. The stability of control-structure interaction is shown to be governed by the ratio of the moments of inertia of the flexible to the rigid portions of the spacecraft. If this ratio is below unity for each axis, the spacecraft is stable; otherwise it is not. The stability inequality for symmetric elastic modes, when they induce attitude motion because of a moment arm from the vehicle mass center, is formulated. Extensive numerical results conclusively establish that the IPFM controllers are eminently suitable for the tasks mentioned above.

Robust Control for Large Space Structures: Spillover Suppression by Frequency-Shaped Optimal Regulator

This paper proposes to give the low-pass property to the LQ regulator in order to enhance the stability robustness against the high-frequency uncertainties. The controller design is formulated as a frequency-shaped regulator problem. Its algorithm and characteristics are studied. The controller is applied to a flexible spacecraft attitude control problem to demonstrate its capability to suppress the spillover instability. Some experimental results are shown.

Near-minimum time open-loop slewing of flexible vehicles

Minimum time, open-loop, optimal controls are calculated for single-axis maneuvers of a flexible structure. By shaping the control profiles with two independent parameters, a wide variety of control histories can be pro- duced. Based on the dynamics of the model, with a normalized time scale, the resulting Pontryagin's necessary conditions yield a nonlinear fixed final time, fixed final state, two-point boundary value problem with the maneuver time as a control parameter. Upon generating numerical solutions to the problem, the final maneuver time and residual flexural energy are compared to the bang-bang solution as a measure of the success of a given maneuver. Examples presented illustrate near-minimum time maneuvers with control of flexible modes in addi- tion to the rigid body modes, as well as the qualitative and quantitative effect of the torque shaping parameters. EAR-MINIMUM time attitude control of flexible spacecraft with active vibration suppression is a topic of current research. Although minimum time/fuel maneuvers have been previously examined,1 such bang-bang controls cause spillover effects that may induce high residual flexural energy. The source of the energy spillover into the flexural modes is the instantaneous switching that results from the bang-bang controls, a potential problem for the control hard- ware (reaction wheels, control moment gyros, etc.) as well. Consequently, minimum-time optimal control problems are often of academic interest only for producing theoretical lower bounds for the maneuver time. We have also found that near bang-bang controls are usually very sensitive to model er- rors; therefore, control shaping is an important issue in ob- taining robust controls. Upon modifying the control profile with a smoothing func- tion and transforming the independent variable (time), a near- minimum-time problem is generated with the mathematical form of a fixed-time nonlinear optimal control problem. The resulting boundary-value problem yields to a number of established methods of numerical solution. The controls are attenuated in such a way that the magnitude of the control rises smoothly from zero to the bounded maximum at the in- itiation of the maneuver and typically has an identically shaped reduction to zero at the final time. In addition, any in- stantaneous switch during the maneuver is shaped using a smooth, continuous function. The sharpness of the control trajectory is determined by a set of arbitrary parameters such that we can produce profiles with low control rates as well as controls that approach, to any desired degree, the bang-bang minimum time solution. The independent variable is transformed such that a linear free-final-time optimal control problem is converted to a nonlinear fixed-time problem, where the maneuver time is contained explicitly as a parameter in the transformed system. This augmented fixed-time problem can then be solved numerically for the controls and the corresponding minimum time required to complete the maneuver.