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This paper addresses the fault-tolerant attitude tracking control problem for spacecraft with limited communication capability and input saturation. The design of a hysteresis quantizer aims to alleviate communication burden between the controller module and the actuator module. Then, the attitude tracking problem for spacecraft with input quantization, input saturation, actuator faults and external disturbances is transformed into an attitude control problem with uncertain input coefficients and bounded disturbances. Thereafter a dynamic loop gain function-based approach and a hyperbolic tangent function term with a time-varying boundary layer are introduced to address the uncertainties of input coefficients and the bounded disturbances, respectively. To facilitate the boundness analysis of the signals in the closed-loop system, a pertinent lemma about the dynamic loop gain function is proved. Finally, numerical simulations are employed to validate the effectiveness of the proposed scheme. • The spacecraft attitude tracking problem with disturbances and uncertainties. • The quantized control law alleviates communication burdens of spacecraft. • A dynamic loop gain function method is used to address the impact of uncertain. • A hyperbolic tangent term is devised to counteract bounded disturbances. [ABSTRACT FROM AUTHOR]

In this paper, straightforward and efficient brute force method is developed to slew spacecraft in the presence of static and dynamic constrained regions. Fuzzy logic is utilized to develop a fuzzy based controller that is used to control the maneuvers and test the proposed method. The controller performance is first tested in an unconstrained attitude maneuvers and then it is utilized to test the proposed method. The spacecraft is assumed to be a 3U CubeSat and the results show through Monte Carlo simulations that the controller has successfully avoided the undesired static and dynamic celestial objects. The proposed method inherits its simplicity from the geometric approach and has the ability to monitor continuously the constraints and it shows that the proposed method works in the presence of multiple constraints. [ABSTRACT FROM AUTHOR]

Copyright of Control Theory & Applications / Kongzhi Lilun Yu Yinyong is the property of Editorial Department of Control Theory & Applications and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Actuator fault poses a challenge to the attitude control of spacecraft. Fault-tolerant control (active or passive) is often used to overcome this challenge. Active methods have better performance than passive methods and can manage a broader range of faults. However, their implementation is more difficult. One reason for this difficulty is the critical reaction time. The system may become unrecoverable if the actual reaction time becomes larger than the critical reaction time. This paper proposes using a feedforward neural network to reduce the actual reaction time in the active fault-tolerant control of spacecraft. Besides this improvement, using a feedforward neural network can increase the success percentage. Success percentage is the ratio of successful simulations to the total number of simulations. Simulation results show that for 200 simulations with random faults and initial conditions, the actual reaction time decreases by 73%, and the success percentage increases by 25%. Based on these results, the proposed controller is a good candidate for practical applications.

Currently, managing a group of satellites or robot manipulators requires coordinating their motion and work in a cooperative way to complete complex tasks. The attitude motion coordination and synchronization problems are challenging since attitude motion evolves in non-Euclidean spaces. Moreover, the equation of motions of the rigid body are highly nonlinear. This paper studies the attitude synchronization problem of a group of fully actuated rigid bodies over a directed communication topology. To design the synchronization control law, we exploit the cascade structure of the rigid body's kinematic and dynamic models. First, we propose a kinematic control law that induces attitude synchronization. As a second step, an angular velocity-tracking control law is designed for the dynamic subsystem. We use the exponential coordinates of rotation to describe the body's attitude. Such coordinates are a natural and minimal parametrization of rotation matrices which almost describe every rotation on the Special Orthogonal group S O (3) . We provide simulation results to show the performance of the proposed synchronization controller. [ABSTRACT FROM AUTHOR]

• A novel linear model-based controller is designed for the CEACS attitude regulation. • The Takagi-Sugeno fuzzy model with a parallel distribution approach has been introduced. • The Fuzzy-MPC controller achieves the desired CEACS attitude pointing accuracy. • The proposed controller effectively handles the CEACS actuator constraints. • The CEACS attitude actuation consumes a lower on-board power. A Combined Energy and Attitude Control System (CEACS) is a dual system in which flywheels are used as energy storage and attitude control devices. This work is a progress on CEACS for small satellites to improve the attitude accuracy. In this maiden work, the Fuzzy–MPC controller is introduced to regulate the CEACS attitude with both linear and nonlinear initial angles, and in the presence of actuator constraints. The nonlinear attitude model is transformed into a set of linear subsystems using the Takagi–Sugeno fuzzy modeling approach. The subsystem is designed with a MPC controller and a total control action is summed up using the parallel distributed compensation approach. The numerical results have been analyzed in terms of attitude pointing accuracies, actuator constraints, and energy consumed by the actuators. A performance comparison between the Fuzzy–MPC and Fuzzy–LQR controllers has also been done. The results validate that the Fuzzy-MPC controller achieves the desired CEACS attitude pointing accuracy of 0. 0010 ° with a zero steady-state error and keeps the control torques within the actuator constraints. On the other hand, the Fuzzy–LQR controller gives some measurable steady-state error and achieves the CEACS attitude pointing accuracy of 0. 0012 °. Therefore, the Fuzzy–MPC based CEACS architecture not only achieves the desired pointing accuracies but also produces ten times smaller control torques than the Fuzzy–LQR controller, which result in a lower onboard power consumption. [ABSTRACT FROM AUTHOR]

• An ESO-based controller is proposed for the attitude control of spacecraft with unknown faults. • The event-triggered policy reduces the effects of constants on the final convergence regions. • The controller reduces the frequency of data transmission and actuators updates. An event-triggered control strategy based on extended state observer (ESO) is proposed for the attitude tracking problem of small plug-and-play spacecraft with uncertain inertia parameters, external disturbances, and actuator faults. A simplified controller is developed based on the angular velocity and the general disturbances estimated by the provided ESO using the information of the system inputs and the angular velocities. In the designed event-triggered sampling mechanism, a state-dependent event-triggered strategy determines the triggering instant of the controller to reduce the frequency of information transmission between the controller and the actuator. In comparison with the previous literature, this paper considers uncertain inertia parameters, external disturbances, and actuator faults as general disturbances estimated by ESO, especially for the actuator faults. The inputs of ESO are the error of the angular velocities, which can simplify the controller design. Moreover, the designed ESO can effectively attenuate the influence of measured noises generated by the gyroscopes. The proposed event-triggered policy balances the performance of event-triggering and the control stability performance, which reduces the final state convergence regions without increasing more triggering times compared to existing studies. Furthermore, the investigated policy achieves Zeno-free triggering. Numerical simulations verify theoretical results. [ABSTRACT FROM AUTHOR]

This article addresses the dynamics and control problem of integrated position and attitude stabilization via manipulator actuation for spacecraft in proximity operations subject to external disturbances. Following recursive modeling philosophy, kinematics, and dynamics are first formulated for multiple-manipulator actuated coupled position and attitude system of spacecraft in the presence of external disturbances, where two vector factorizations are proposed to ensure a compact and explicit dynamic formulation, holding the inherent skew-symmetric property of the coefficient matrices. In view of the unconspicuous cascaded-like system structure, a nonlinear control scheme is designed following backstepping philosophy to drive the joints of manipulators causing reactions to robustly stabilize the spacecraft position and attitude. Toward this end, a reference trajectory prescribing the spacecraft motion is predesigned by means of polynomial functions ensuring well-behaviored performance. Then, as a key component, a second-order dynamic filter is constructed making use of the dynamics structure to generate the command joint motion capable of manipulator actuation. The rigorous closed-loop stability and robust performance analyses are undertaken within the Lyapunov framework. Moreover, three types of self-collisions possibly happening in the motion of multiple manipulators are considered and handled by three collision detection methods, which facilitates a collision-free trajectory shaping during the aforementioned reference trajectory design phase. Finally, numerical simulations are given to demonstrate the effect and robustness of the proposed control scheme. [ABSTRACT FROM AUTHOR]

The predefined-time tracking problem of spacecraft attitude systems with loss of actuator effectiveness and lumped disturbance including the external disturbance and system uncertainty is studied. In order to obtain the estimation of the actuator efficiency factor more quickly and accurately, a robust learning observer is designed. Based on the fault reconstruction information of the learning observer and the predefined-time stability lemma, a predefined-time tracking fault-tolerant control scheme is proposed for the faulty spacecraft attitude system. The stability of the learning observer and the whole control system is verified by the Lyapunov stability theory. Finally, the effectiveness and advantages of the proposed scheme are illustrated by simulation results including comparisons with existing works. [ABSTRACT FROM AUTHOR]

As large-sized spacecraft have been developed, they have been equipped with flexible appendages, such as solar cell plates and mechanical flexible arms. The attitude control of spacecraft with flexible appendages has become more complex, with higher requirements. In this paper, a fractional-order PD attitude control method for a type of spacecraft with flexible appendages is presented. Firstly, a lumped parameter model of a spacecraft with flexible appendages is constructed, which provides the transfer function of the attitude angle and external moment. Then, a design method for the fractional-order PD controller for the attitude control of a spacecraft with flexible appendages is provided. Based on the designed steps, a numerical example is provided to compare the control performances between the fractional-order and integer-order PD controllers. Finally, the obtained numerical results are presented to verify the effectiveness of the proposed control method. [ABSTRACT FROM AUTHOR]

This paper is concerned with the agile rest-to-rest attitude maneuvering of spacecraft using pyramid-type single-gimbal control moment gyroscopes (SGCMGs). In this method, the gimbal rate command to the SGCMGs is determined by calculating the pair of optimal trajectories of the attitude and the gimbal angles without any "steering law". Thus, the proposed method does not suffer from the singularity problem causing the increase in control error and settling time. Before maneuvering, the pair of the initial optimal trajectories of the attitude and the gimbal angles is calculated by the minimum-time problem. Then, to compensate for errors in maneuvering caused by disturbances and/or modeling errors, the trajectories of the attitude and the gimbal angles are recalculated and updated frequently during the control. In this paper, optimization problems for generating and updating the trajectories are formulated, and specific control algorithms are provided. Finally, numerical simulations and ground experiments with an air floating table are performed to verify the feasibility and effectiveness of the proposed control method. For the rest-to-rest attitude maneuver, the proposed method can be adapted not only to the presence of modeling error in the inertia tensor, but also to the failure of one or two CMGs in the SGCMG system, without changing the control algorithm. These demonstrate the high robustness and practicality of the proposed method. [ABSTRACT FROM AUTHOR]

Currently, managing a group of satellites or robot manipulators requires coordinating their motion and work in a cooperative way to complete complex tasks. The attitude motion coordination and synchronization problems are challenging since attitude motion evolves in non-Euclidean spaces. Moreover, the equation of motions of the rigid body are highly nonlinear. This paper studies the attitude synchronization problem of a group of fully actuated rigid bodies over a directed communication topology. To design the synchronization control law, we exploit the cascade structure of the rigid body’s kinematic and dynamic models. First, we propose a kinematic control law that induces attitude synchronization. As a second step, an angular velocity-tracking control law is designed for the dynamic subsystem. We use the exponential coordinates of rotation to describe the body’s attitude. Such coordinates are a natural and minimal parametrization of rotation matrices which almost describe every rotation on the Special Orthogonal group SO(3). We provide simulation results to show the performance of the proposed synchronization controller.

The problem of developing a controller for spacecraft to accomplish attitude maneuvers within a prescribed time, independently of the initial states, is addressed. A fixed-time disturbance observer-based control is established and proved to drive the error state to zero, in a prescribed time, in the presence of modeling and external uncertainties. The key advantages of the proposed control are that the gains are explicitly determined by the prescribed maneuvering time only and that it is also robust to instantaneous disturbances. The control approach enables an efficient design process for attitude control applications with time constraints. Simulation results are presented for the attitude control of a rigid spacecraft to verify the benefits of this control scheme. [ABSTRACT FROM AUTHOR]

In this article, the problem of optimal time-varying attitude tracking control for rigid spacecraft with system constraints and unknown additive disturbances is considered. Through the design of a new nonlinear tube-based robust model predictive control (TRMPC) algorithm, a dual-loop cascaded tracking control framework is established. The proposed TRMPC algorithm explicitly considers the effect of disturbances and applies tightened system constraints to predict the motion of the nominal system. The obtained optimal control action is then combined with a nonlinear feedback law such that the actual system trajectories can always be steered within a tube region centered around the nominal solution. To facilitate the recursive feasibility of the optimization process and guarantee the input-to-state stability of the tracking control process, the terminal controller and the corresponding terminal invariant set are also constructed. The effectiveness of using the proposed dual-loop TRMPC control scheme to track reference attitude trajectories is validated by experimental studies. A number of comparative studies were carried out, and the obtained results reveal that the proposed design is able to achieve more promising constraint handling and attitude tracking performance than that of the other newly developed methods investigated in this research. [ABSTRACT FROM AUTHOR]

This article investigates the distributed adaptive attitude synchronization problem for multiple rigid spacecraft with unknown inertial matrices and event-triggered communication. Under a directed topology condition, two distributed event-based adaptive attitude synchronization schemes are proposed. By introducing a reference system for each spacecraft, a distributed adaptive control algorithm is first developed, based on which a group of rigid spacecraft can be synchronized at a common attitude with zero angular velocity. Then, the case without angular velocity measurements is considered. By redesigning the reference system and resorting to a pseudovelocity filter, an event-based adaptive output feedback attitude synchronization algorithm is presented. It is shown that with the proposed control schemes, all the closed-loop signals are uniformly bounded and the attitude synchronization can be achieved, while Zeno behavior in each spacecraft can be excluded. Simulation results are provided to illustrate the effectiveness of the proposed control schemes. [ABSTRACT FROM AUTHOR]

This article focuses on the problem of finite‐time attitude tracking control for a rigid spacecraft with event‐triggered inputs. The existing event‐based attitude control methods are developed from the smooth control law and can only provide at best Lipschitzian exponentially convergent. Because the nonsmooth control law shows a faster convergent rate and better control performance, a new event‐triggered nonsmooth attitude tracking controller is proposed through the emulation method of a continuous‐time finite‐time attitude controller. With the aid of Lyapunov theory and finite‐time stability lemma, the stability analysis of the closed‐loop system is given and the event‐triggered control system is Zeno‐free. Like the case of a continuous‐time nonsmooth attitude controller, it is proved that the event‐triggered nonsmooth attitude controller maintains a better disturbance rejection ability. Finally, some simulation results are given to verify the theoretical results. [ABSTRACT FROM AUTHOR]

This study surveys the developments in satellite attitude determination and control system, especially for microsats. This survey is not intended to be complete but is limited to the most significant developments of sensors, actuators, and algorithms in the last two decades. First, attitude determination methods including algorithms and sensors together with actuator-based control methods are introduced. Furthermore, current problems in alignment error, flexible satellites, and low redundancy of microsats attitude determination and control system are discussed. Moreover, developments of some deep-neural-networks-based methods, which have great potential in solving current problems, are summarized. [ABSTRACT FROM AUTHOR]

The algorithm of the turn of the spacecraft from an initial arbitrary angular position at an arbitrary angular rate to a solar-oriented attitude is investigated. Minimum essential equipment of the motion control system required for the purpose of ensuring maintenance of solar orientation is defined: a solar sensor, a single-axis angular-rate sensor, low-thrust liquid rocket engines. A solution of the problem of defining the spacecraft angular rate vector by the measurements of the deviation of the optical axis of the solar sensor from the sun vector and the single-axis angular-rate sensor is presented. The conditions under which control action on the rocket engines for the purpose of changing the value of the angular-rate vector for the Sun to get into the field of viewing of the solar sensor or for the spacecraft stabilization are defined. Mathematical modeling of the spacecraft attitude control system with the unknown initial state vector of motion is carried out. The results of mathematical modeling confirmed the efficiency of the proposed algorithm in terms of reducing propellant fuel consumption and high-speed performance. In comparison with the known methods of solving the problem of reducing angular speed (a lengthy process with the use of a magnetic system or a fast process with the use of a three-axis angular-rate sensor and rocket engines) the duration of the process of reducing angular speeds is the same as in normal operation, however, at the same time the problem of bringing spacecraft into the solar oriented attitude is solved.

As large-sized spacecraft have been developed, they have been equipped with flexible appendages, such as solar cell plates and mechanical flexible arms. The attitude control of spacecraft with flexible appendages has become more complex, with higher requirements. In this paper, a fractional-order PD attitude control method for a type of spacecraft with flexible appendages is presented. Firstly, a lumped parameter model of a spacecraft with flexible appendages is constructed, which provides the transfer function of the attitude angle and external moment. Then, a design method for the fractional-order PD controller for the attitude control of a spacecraft with flexible appendages is provided. Based on the designed steps, a numerical example is provided to compare the control performances between the fractional-order and integer-order PD controllers. Finally, the obtained numerical results are presented to verify the effectiveness of the proposed control method.

The attitude dynamics of a nanosatellite (NS) with one movable unit changing its angular position relative to a main body of the nanosatellite is considered. This relative movability of the unit can be implemented with the help of flexible rods of variable length connecting the unit with the main body. Change of the relative position of the movable unit shifts the center of mass of the entire mechanical system. The NS has a single jet engine rigidly mounted into the NS main body. The shift of the mass center creates an arm of the jet-engine thrust and a corresponding control torque. Schemes to control the attitude dynamics of the satellite using the movability of its unit are developed, using both the torque from the engine and inertia change. [ABSTRACT FROM AUTHOR]

For CubeSats requiring high pointing accuracy and slewing agility, ground-based hardware-in-the-loop simulations are strongly demanded to test and validate spacecraft subsystems and guidance, navigation, and control algorithms. In this article, a magnetic attitude control system (MACS) testbed for a CubeSat is developed utilizing a spherical air bearing and a Helmholtz cage. The design, development, and verification procedure of MACS is presented together with different test scenarios. To generate enough torque with the magnetorquer system in the dynamic testbed, the Helmholtz coil system of the testbed has driven to provide an augmented magnetic field. As an example of experimentation, the B-dot control algorithm was implemented to dissipate the angular momentum of the dynamic MACS testbed. The experimental results were compared with those of the numerical simulations. [ABSTRACT FROM AUTHOR]

In this paper, an adaptive nonsingular terminal sliding mode control (ANTSMC) is investigated for attitude tracking of spacecraft with actuator faults. First, a nonsingular fast terminal sliding mode surface is designed to avoid the singularity. Finite-time attitude control is developed using the nonsingular terminal sliding mode technique, which can make the attitude and angular velocity tracking errors converge to zero in finite time in the presence of uncertainties and external disturbances. Second, the total uncertainty is deduced to be bounded. The adaptive laws are incorporated to develop the ANTSMC, removing the restriction on the upper bound of the lumped uncertainty. The finite-time convergence of the closed-loop system with ANTSMC is proved using the Lyapunov stability theory. Finally, the simulation results are presented to demonstrate the performance of the proposed controllers.

The attitude stabilization of flexible spacecraft is significantly affected by environmental disturbances, elastic vibration of flexible appendages, and model uncertainties. These disturbances and uncertainties can be accurately estimated by the extended disturbance observer (EDO). However, the EDO is sensitive to measurement noise when high-order and large bandwidth are chosen. Thus, the output of the EDO will contain too much high-frequency noise which can degrade the attitude stability of the spacecraft and cause heavier chattering of the actuators. In order to suppress the effect of measurement noise, a noise reduction EDO (NREDO) is proposed in this paper. First, a dynamical model of disturbance is reconstructed by augmenting the integral of the virtual measurement of the lumped disturbance as a new state. Second, the NREDO is designed by utilizing the estimation error of the augmented state to update the observer estimates. Thereafter, the disturbance estimation accuracy is analyzed for the EDO and NREDO in the frequency domain in the presence of measurement noises. Finally, the simulations are conducted to verify the effectiveness of the proposed controller. The theoretical analysis and simulation results demonstrated that the NREDO can not only effectively suppress the effect of measurement noise, but also inherit the advantages of the EDO in avoiding redundant estimation of the system states and dealing with the nonlinear item of the spacecraft model. Improved attitude control performance can be achieved under the proposed NREDO-based controller.

In this article, a novel distributed attitude coordination control scheme is proposed for a group of rigid spacecraft with unknown disturbances. The communication topology is described by a general directed graph, and the leader's attitude is available to only a small fraction of followers. Different from existing attitude coordination control schemes, here an event-triggered communication strategy without Zeno phenomenon is designed, which avoids continuous communication between neighboring spacecraft and considerably reduces the communication burden. Moreover, a nonlinear observer is constructed for each follower, which enables us to estimate and compensate for a large class of disturbances exactly. The proposed scheme is able to ensure the boundedness of all closed-loop signals and steer the attitude tracking errors into an arbitrarily small residual set, regardless of the intermittent communication. Simulation results illustrate the effectiveness of the proposed scheme. [ABSTRACT FROM AUTHOR]

In this paper, we develop a new gain-scheduled (GS) controller for attitude control of spacecraft when the plant models are given as linear parameter-varying (LPV) systems. Avoiding trial and error in determining the weighting matrices of the H2 performance evaluation, we introduce the inverse linear quadratic (ILQ) method, while generalizing the existing theory for linear time-invariant (LTI) systems to that for LPV systems. The existing optimality condition is also rebuilt in a linear matrix inequality (LMI) condition and generalized for LPV systems. As a numerical example, in this paper, we deal with angular velocity control of a spacecraft with reaction wheels (RWs) and demonstrate the effectiveness of the proposed controller through a numerical simulation [ABSTRACT FROM AUTHOR]

Copyright of Control Theory & Applications / Kongzhi Lilun Yu Yinyong is the property of Editorial Department of Control Theory & Applications and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

In this paper, the attitude control of spacecraft using a pyramid-type variable-speed control moment gyro (VSCMG) system is considered. For successive attitude maneuvers, it is desirable that the wheel angular momentums are regulated to nominal values after an attitude maneuver. In addition, avoidance of reference aligned singularity (RAS) is also important. RAS is a singularity point where the singular direction of the constant-speed control moment gyro (CSCMG) system is aligned to a reference torque. Regarding the motivations, this paper proposes a steering law which achieves attitude control and regulates the wheel angular momentums while successfully avoiding RAS. Both numerical simulations and experimental verification are performed to demonstrate not only the validity of the proposed method, but also a certain robustness against unmodeled dynamics and the gravity torque. • Spacecraft attitude is controlled by variable speed control moment gyros. • Reference aligned singularities of CMGs are particularly focused in the control. • Wheel angular momentums are usually regulated to nominal values. • Effectiveness of the control law is validated in all maneuver directions. • Ground experiments are conducted in order to examine the proposed control. [ABSTRACT FROM AUTHOR]

With the rapid development of micro-spacecraft, miniaturization has become a key problem in attitude control. As an angular momentum exchange actuator, a single reaction sphere can generate multi-DOF output torques and satisfy the requirement of three-axis attitude control. This enables it to be a potential alternative for micro-spacecraft attitude control system. Categorized by the driving principles, the various design techniques of reaction spheres are firstly reviewed, including the spherical motor based and the multi-driving-wheels based. Their advantages and disadvantages are summarized. In order to realize the closed-loop control of the reaction sphere, the techniques of measuring the speed and orientation of the reaction spheres are subsequently reviewed. These include rotary encoders or gyroscope with gimbal guideway, photoelectric encoding ring, machine vision, Hall sensor, color sensor and piezoelectric sensor. By promoting the popularization of the concept of reaction spheres, it is expected to provide new ideas for the miniaturization design and attitude control of micro-spacecraft. • The current designs of reaction spheres are reviewed, whose driving methods are categorized into spherical motors and multi-driving-wheels. • The speed and orientation measurement methods of the reaction spheres are reviewed, which are essential for their real-time control. • Permanent magnet reaction spheres have advantages of miniaturization and low energy consumptions among all the driving methods, combined with Hall sensors and optical mouse sensors for feedback control. [ABSTRACT FROM AUTHOR]

This paper is devoted to attitude tracking control of fractionated spacecraft with wireless communication. We consider the practical case that the spacecraft suffers from uncertain inertia parameters, external disturbances, and even unknown and time-varying actuator faults. Within the framework of the backstepping method, a novel event-triggered adaptive fault-tolerant control scheme is proposed. In our design, an event-triggering mechanism is introduced to determine the time instants for communication, which successfully avoids continuous communication and Zeno phenomenon. Then, with the aid of a bound estimation approach and a smooth function, the impacts of the actuator faults, as well as the network-induced error, are effectively compensated for. Moreover, by employing the prescribed performance control technique, it is shown that the attitude tracking errors can converge to predefined arbitrarily small residual sets with prescribed convergence rate and maximum overshoot, no matter if there exist unknown actuator faults. Compared with conventional adaptive attitude control schemes, the proposed scheme significantly reduces the communication burden, while providing high reliability and stable, rapid, and accurate response for attitude maneuvers. Simulation results are presented to illustrate the effectiveness of the proposed scheme. [ABSTRACT FROM AUTHOR]

This brief examines the problem of attitude tracking control with prescribed performance guarantees for a spacecraft subjected to actuator faults and input saturation. To pursue this, the open-loop tracking error dynamics with certain designer-specified performance constraints is first transformed into an equivalent “state-constrained” one, via an error transformation; furthermore, the resulting dynamics is augmented with a dynamic system, which is tactfully constructed to ensure that the control input satisfies the magnitude limits. Subsequently, a robust fault-tolerant controller is developed by using a low-pass filter and an auxiliary system in conjunction with adaptive backstepping design. It is shown that the control algorithm developed not only achieves the stable attitude tracking with prescribed behavioral metrics but also guarantees the boundedness of all the closed-loop signals. Finally, simulation results are given to evaluate the efficacy of the proposed scheme. [ABSTRACT FROM AUTHOR]

The spacecraft attitude tracking problem with exponential convergence and resilient control capability ensured is investigated. The system uncertainties induced by uncertain inertia and external disturbance, and the actuator uncertainties consisting of misalignment and faults, are addressed simultaneously. An observer-based exponential and resilient control scheme is presented. The closed-loop system is ensured to be stable that the attitude and the angular velocity tracking errors converge with an exponential rate to a small set with an arbitrary radius. The fast and high accuracy attitude tracking maneuver is accomplished. This resilient controller, characterized by a simple structure and less onboard computation, has great practical application potential. A rigid spacecraft example is shown to verify the effectiveness of the resilient solution with simulation and experimental results presented. [ABSTRACT FROM AUTHOR]

In this study a decentralised event‐based scheme is proposed to synchronise the spacecraft's attitudes during their convergence toward the desired attitude, in the presence of disturbance. To achieve this goal, decentralised integral‐based triggering mechanisms are introduced which not only reduce the data transmission among the spacecraft efficiently but also guarantee a minimum inter‐event time for each triggering module. Moreover, a design procedure is given to find the control gains and triggering conditions parameters properly. Finally, simulation results are provided to show the effectiveness of the system. [ABSTRACT FROM AUTHOR]

In this paper, we establish the dynamics of a spacecraft with Variable-Speed Control Moment Gyros (VSCMGs). Based on this dynamics, we develop an easy-to-use Linear Parameter-Varying (LPV) model, in which we deal with 3-axis attitude control of a spacecraft with pyramid-array VSCMGs. Although it is hard to consider overall stability and control performance of attitude control at the same time with conventional Lyapunov function-based controllers, this paper attains both of them via LPV control theory. And also we design two types of Gain-Scheduled (GS) steering laws for singularity avoidance and escape. Through numerical examples, the efficiency of the proposed GS controller and the GS steering laws is demonstrated.

This paper addresses the difficult problem of fault-tolerant control and closed-loop control allocation for spacecraft attitude control system with actuator failures, actuator saturation, and external disturbances. As a fundamental step, a modified fault diagnosis observer is proposed utilizing the iterative learning methodology to reconstruct the actuator failures in real time. On the basis of the previously reconstructed failure information, a fault-tolerant control scheme incorporating with a parameter adjusting law is presented to enforce the spacecraft attitude control system to reach the real sliding mode surface in finite time. Meanwhile, the singularity of control command will be avoided using the saturation function, and the chattering problem restrained by the adjusting law. Furthermore, with the concept of control allocation, a novel on-line closed-loop constrained optimal fault-tolerant control allocation scheme is employed to distribute the signals synthesized by the baseline controller over the redundant actuators with failures and constraints. The stability of the closed-loop fault-tolerant control allocation process is guaranteed by the theory of the closed-loop discrete-time feedback control system. The key achievement of the proposed systematic strategy is that the whole closed-loop attitude control system can theoretically be guaranteed to be stable. Numerical simulations are carried out to verify the effectiveness and superiority of the developed approach. [ABSTRACT FROM AUTHOR]

The postcapture of a partially failed spacecraft by space manipulators will bring a mutation in the dynamics of the combination. Both the inertia properties and the thruster configuration matrix will change significantly. The unknown dynamics of the partially failed spacecraft also cause a tremendous technical challenge for attitude takeover control. Accordingly, this paper describes a novel reconfigurable control system for postcapture attitude takeover of a partially failed spacecraft with parametric uncertainties, whose fuel has been exhausted or whose actuators have partial malfunctions. First, the reconfigurable control law is designed by command filtering adaptive back-stepping control to guarantee the system performance and global asymptotic stability considering inertia parametric uncertainties. Second, the thrusters are reconstituted, without changing the thruster physical configuration. Finally, the thrusters’ forces can be redistributed by the dynamic control reallocation method based on constrained quadratic programing. Numerical simulations validate the feasibility of the proposed approach for postcapture attitude takeover control of a partially failed spacecraft with parametric uncertainties. Note to Practitioners—This paper presents a methodology for a partially failed spacecraft with parametric uncertainties. It focuses on solving the attitude takeover control of the partially failed spacecraft perfectly, while considering the position and speed constraints of the actuator and avoiding the plume impact to the spacecraft. The proposed command filtering adaptive back-stepping control can be used for spacecraft’s thruster reconfiguration considering the parametric uncertainties. Furthermore, the dynamic control reallocation method is particularly useful for future applications that include spacecraft with redundant actuators. [ABSTRACT FROM AUTHOR]

A passivity-based controller with quantization for spacecraft attitude control is developed. This passive control scheme includes two parts which are a proportional controller for quaternion feedback and a strictly positive real controller for the angular velocity. To alleviate the errors caused by quantization, a special modification for the nonlinear quantized input is employed in the strictly positive real controller. Asymptotic stability can be guaranteed with the presented controller structure. A guideline for the controller parameter selection is provided with sensitivity analysis for the control scheme. Numerical simulation results demonstrate the effectiveness of the proposed controller. [ABSTRACT FROM AUTHOR]

Copyright of Journal of Harbin Institute of Technology. Social Sciences Edition / Haerbin Gongye Daxue Xuebao. Shehui Kexue Ban is the property of Harbin Institute of Technology and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Copyright of Transactions of Beijing Institute of Technology is the property of Beijing University of Technology and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

A quaternion-based attitude unscented Kalman filter is formulated with quaternion errors parameterized by small angle approximations and is applied to a filter with a state vector consisting of the attitude quaternion and the gyro bias vector. The filter is evaluated using extensive Monte Carlo data in a simulated lost-in-space scenario of a low-Earth orbiting spacecraft processing only three-axis magnetometer and gyro measurements. The filter is found to be robust, accurate, and rapidly convergent in this scenario for small true gyro biases and small initial uncertainties in their values, often converging in only one half of an orbit period to an attitude accuracy of 0.1 degrees. The filter convergence is found to depend significantly on the value of the true gyro biases as well as the initial gyro bias covariances. Monte Carlo results also indicate that this unscented Kalman filter is significantly less robust than an extended Kalman filter with the same attitude approach, but performs slightly better than another unscented Kalman filter with a generalized Rodrigues parameter approach to quaternion errors.