This paper designs an active fault-tolerant control system for spacecraft attitude control in the presence of actuator faults, fault estimation errors, and control input constraints. The developed fault-tolerant control system is able to detect the actuator fault without false alarms caused by external disturbances, and also estimate the total fault effects accurately through an indirect fault identification approach, in which an auxiliary variable is utilized to build the relation between fault and system states. Once the fault identification is completed with certain degree of reconstruction accuracy, a fault-tolerant backstepping controller using the nonlinear virtual control input is reconfigured to accommodate the detected actuator faults effectively, in spite of actuator saturation limitations and fault estimation errors. Numerical simulation is carried out to demonstrate that the proposed active fault-tolerant control system is successful in fault detection, identification, and controller reconfiguration for handling actuator faults in attitude control systems.

Aiming at solving the model-free fault-tolerant spacecraft attitude control problem, a data-driven adaptive control scheme is proposed to the spacecraft in the presence of actuator faults and performance constraints. First, the discrete-time attitude dynamic model is transformed into an affine nonlinear system based on local dynamic linearization. Then, a fuzzy logic-based lazily adapted constant kinky inference rule is introduced to predict the arbitrarily continuous nonlinear actuator faults and model uncertainties by supervised learning with insufficient prior knowledge. To satisfy time-varying deferred asymmetric constraints of the attitude tracking error, a virtual control law is proposed in the attitude control loop using the back-stepping approach, which is derived from a deferred switching transformation and barrier Lyapunov function. The stability of the data-driven adaptive fault-tolerant attitude control for the nonlinear discrete-time spacecraft attitude dynamics is analyzed rigorously with the aid of contraction mapping principle and discrete-time Lyapunov theory. Compared with existing methods, the proposed one considers nominal non-global Lipschitz nonlinear system with arbitrarily continuous unmodeled uncertainties and time-varying actuator faults, and achieves smaller prediction error bound than general kinky inference scheme and better closed-loop performance by estimating and compensating unknown dynamics. Finally, numerical simulation verifies the effectiveness of the proposed control scheme.

This brief addresses the single-gimbal control moment gyro (SGCMG) fault modeling, estimation, and tolerant-control steering logic design problem, aiming at enhancing the reliability and safety of spacecraft attitude control systems. The SGCMG is modeled as a two-loop system, including a wheel speed control loop and a gimbal rate control loop. Each loop contains an electrical motor (EM) and its corresponding variable speed drive (VSD), which may suffer from faults. By analyzing and modeling potential faults of the EM-VSD system, the SGCMG fault model is further developed. Then, a local adaptive fault estimator is proposed to reconstruct the total time-varying fault effects of each SGCMG. It is proven that the gimbal angle estimation error and fault estimation error converge to small compact sets containing zero. Moreover, leveraging estimated fault effects, a fault-tolerant steering logic is further developed to allocate the commanded attitude control torque properly such that the gimbal rate constraints are satisfied, and fault effects are compensated. To verify the proposed fault estimator and fault-tolerant steering logic, numerical simulations are carried out on an SGCMG-actuated spacecraft.

This correspondence investigates the under-actuated spacecraft attitude maneuver problem using two parallel single gimbal control moment gyros (SGCMGs). According to the configuration of the SGCMGs, the principle axes of the spacecraft are divided into two parts: namely the axes with and without direct control input. Based on this division, a two-level control strategy: a low level tracking control part to track the desired angular velocity error about the axes with direct control input and a high level sliding mode control part to stabilize the angular velocity error, is developed. Four numerical scenarios are conducted to verify the proposed control logic. Simulation results show that the target attitude is a global equilibrium, and any attitude can be reached when the whole system is feasible. If the controllability is not satisfied, the spacecraft rotates in the adjacent region around the desired attitude.

In this paper, the constrained attitude control prob-lem with a prescribed performance requirement is investigated, and a systematic method involving the synthesis of path planning, prescribed performance design, and tracking control with inherent saturation is proposed. The overall works are based on attitude rotational matrix and its equivalent vector description. In the path planning phase, both the solid constraints, referring to the forbidden region where the sensitive instrument is strictly prohibited, and the soft constraints, referring to a designable margin of safety are considered. Then a centric reference trajectory outside the soft constraints is established according to a gradient-base algorithm. To formulate the reference trajectory, a motion planning method is employed to establish the ideal angular velocity constrained by the maximum affordable control torque. In the prescribed performance design phase, the performance boundary tube is generated to avoid any potential invasion caused by the attitude tracking error. Finally, in the controller design phase, inherent input saturation is considered and a magnitude-constrained controller is proposed. When the proposed controller is exerted on the system, closed-loop system stability is achieved and the spacecraft is always kept outside the solid constraints. The effectiveness of the proposed method is verified through numerical simulations.