جستجوی مقالات مرتبط با کلیدواژه "flexible spacecraft" در نشریات گروه "مکانیک"

In this paper, a robust adaptive hybrid control approach based on a combination of super-twisting and non-singular terminal sliding mode control (STNSMC) approaches for vibration and attitude control of a flexible spacecraft with fully coupled dynamic is developed. The proposed adaptation law eliminates the need for bounds knowledge of external disturbances and uncertainties. Then an ST-based NSMC generates a continuous control signal to reject the Chattering phenomenon, the non-singular terminal switching control law with the ability to generate continuous control commands to eliminate the chattering phenomenon. Moreover, finite-time convergence is achieved, and the singularity problem has been avoided. The overall stability of the system has been demonstrated using the Lyapunov theory. One of the essential features of the proposed control algorithm is to prevent overestimation of control gains and faster convergence rates comparing to conventional ST and non-singular terminal SMC approaches. The simulations in the form of a comparative study for large-angle maneuver reveal the advantage of the proposed approach.

Maneuvering and controlling the attitude with the highest accuracy, speed and lowest power consumption has always been one of the challenges in the field of spacecraft control system design. The flexibility of satellites causes a change in the dynamics of the whole system. In this article, the Robust Quasi-Sliding Mode Control (RQSMC) method has been utilized to control the attitude of the flexible spacecraft in the presence of disturbances. Considering the non-linearity of the equations of the flexible spacecraft, the impossibility of modeling this system ideally and the inability of mathematical descriptions to fully explain the movement of this flight system, this article uses the RQSMC method as a suitable idea for the attitude control of the flexible spacecraft. For this purpose, the three-degree-of-freedom model of the flexible spacecraft including different disturbances in each direction under quaternion kinematics will be presented in the dynamic modeling section of this article, and then a RQSMC will be designed that also has the ability to control chattering.  The checking of functional parameters such as energy consumption index, agility index and body angular rates constraints showed that this controller has a desirable performance in accurate and fast spacecraft orientation.

An active vibration control algorithm and robust integral sliding mode control (SMC) are discussed to stabilize the attitude of the flexible spacecraft under external disturbances and actuator faults. As a coupled rigid-flexible dynamical system, the flexible spacecraft is modeled as a rigid hub with two solar panels equipped with piezoelectric (PZT) sensors and actuators. A passive fault-tolerant integral sliding mode control algorithm using a nominal proportional-derivative control algorithm and an improved fault-tolerant algorithm with time-varying additive fault is developed to increase system’s performance, prevent the system's flexible modes excitations in the phase of reaching the sliding surface. Therefore, when the system enters the sliding mode, the closed-loop dynamic behavior, including actuator faults, will be identical to that of the system without faults. It is possible to reduce the residual vibrations caused by the attitude dynamics and actuator faults by simultaneously activating the strain rate feedback (SRF) vibration control algorithm during the maneuver. The performance of the proposed integral fault-tolerant control in terms of the flexible modes excitation, the control effort, and achieving the desired attitude parameters in a comparative study demonstrated its advantage and superiority over the conventional integral sliding mode algorithms.

This paper presents a study concerning active vibration control of a smart flexible spacecraft during attitude maneuver using thrusters and reaction wheels (RW) in combination and piezoelectric (PZT) sensor/actuator patches. The large-angle maneuver and residual vibration of the spacecraft are controlled using an extended Lyapunov-based design (ELD) and strain rate feedback (SRF) theory for a two-mode mission. The single-axis fully coupled nolinear rigid-flexible dynamic of the system is derived applying a Lagrangian approach and Finite Element Method (FEM). The overall stability of the system including energetic terms covering a hub and two flexible appendages, torsional spring, RW, and PZT dynamics, has been proved and the control law has been derived accordingly. A pulse-width pulse-frequency (PWPF) modulation is used to alleviate the excitations of high-frequency flexible modes. However, due to the fast maneuver, there are still residual vibrations in the system. Hence, the SRF algorithm using PZT is applied to prepare further vibration suppression. A great feature of the proposed hybrid actuator system is the switching time of the thrusters and RW, which is based on total system energy. The numerical results for a flexible spacecraft with large-angle, agile, and precise maneuver requirements through a comparative study verify the merits of the proposed approach.

This paper deals with active vibration control (AVC) and structural health monitoring of a maneuvering flexible spacecraft with a cracked panel using piezoelectric (PZT) sensor/actuator patches and strain rate feedback (SRF) method. The cracked panel is modeled using the Euler-Bernoulli beam theory and the finite element method (FEM). The nonlinear equations of motion of the fully coupled rigid-flexible system are extracted using the Lagrangian formulation and solved with Newmark-β algorithm. Two approaches of structural health monitoring; first, trial and error, and second, maximum strain rates (online measurement) along with AVC (applying control signals based on the maximum values of strain rates to the predefined number, but unknown locations of PZT actuators), are performed. The strain rates change directly with mission conditions and crack locations, and the corresponding actuators are activated simultaneously. Moreover, to identify the dynamic behavior of the whole, cracked system, an energy function with different weighting coefficients is introduced to propose a suitable criterion for the performance evaluation of the second approach. Simulations for different crack numbers and locations and external disturbances as a comparative study (in MATLAB/Simulink) demonstrate suitable criteria to determine the number and locations of actuators and reduce the power costs in modern spacecraft in high-precision missions.

This paper presents a study concerning active vibration control of a smart, flexible spacecraft during attitude maneuver using thrusters/ a reaction wheel and piezoelectric patches. The large-angle maneuver and residual vibration of the spacecraft are controlled utilizing an extended Lyapunov-based design (ELD) and strain rate feedback (SRF) theory. The single-axis fully coupled rigid-flexible dynamic of the system is derived applying a Lagrangian approach and Assumed Mode Method (AMM). The system's overall stability, including energetic terms covering a hub, two flexible appendages, PZT sensor/actuator, RW dynamics, and torsional spring, has been proved, and the control law has been derived accordingly. A pulse-width pulse-frequency (PWPF) modulation is used to alleviate the excitations of high-frequency flexible modes. However, due to the fast maneuver, there are still residual vibrations in the system. Hence, the SRF algorithm using PZT is applied to prepare further vibration suppression. The performance of the proposed extended controller is compared to the conventional Lyapunov and pole placement control algorithms. The numerical results for simultaneously large angle attitude and vibration control of a flexible spacecraft through a comparative study verify the merits of the proposed approach.

This paper analyses the dynamic behavior of the rigid solar panels deploying mechanism of a spacecraft with flexible hinges. The proposed mechanism, maintaining a proper speed, guarantees the deployment synchronization of solar panels and minimizes the effects of impact and vibration applied during the final stage and after the panels lock-up using torsional springs in the hinges and yoke driven assembly. The equations of the motion of the system are derived using Lagrangian approach and the behavior of the mechanism for constant and variable torque excitation modes is investigated. The simulation results presented along with the dynamic simulations performed by Adams software and conventional mechanisms show the efficiency of the proposed method.

This paper presents advanced controllers for maneuver and vibration control of dynamic systems with rigid-flexible coordinates. Stability analysis of such a control system is one of the important issues. The control of flexible spacecrafts or flexible manipulators with rigid-flexible body dynamics has challenges that encourage designers to develop advanced controllers. Each of these control systems has some advantages and some disadvantages. The purpose of the present paper is to study various control design methods and to review the literature by addressing the weaknesses and strengths of existing techniques.

In this paper, equation of motion of three axis attitude dynamic of flexible spacecraft is derived using combination of finite element method and Euler equation. Flexible appendafes are modeled by beam elements. Goal of control is target attitude of spacecraft from initial state to desired attitude and suppression of vibration that induced in flexible appendages. So a combination of backstepping and sliding mode control method used for three-axis attitude maneuver of flexible spacecraft and for suppressing vibration of flexible appendage used from active vibration control method by PZT actuator. Control law for vibration control is based on LQG method.