Search results

Recent spacecraft attitude control systems tend to use wireless communication for cost-saving and distributed mission purposes while encountering limited communication resources and data exposure issues. This paper aims to study the attitude control problem with low communication frequency under the sampled-data.

The authors propose constructive control system structures based on quantization and event-triggered methods for intra-spacecraft and multi-spacecraft systems, and they also provide potential solutions to shield the control system's data security. The proposed control architectures can effectively save communication resources for both intra-spacecraft and multi-spacecraft systems.

The proposed control architectures no longer require sensors with trigger-ing mechanism and can achieve distributed control schemes. This paper also provides proposals of employing the public key encryption to secure the data in control-loop, which is transmitted by the event-triggered control mechanism.

Spacecraft attempts to use wireless communication, yet the attitude control system does not follow up promptly to accommodate these variations. Compared with existing approaches, the proposed control structures can save communication resources of control-loop in multi-sections effectively, and systematically, by rationally configuring the location of quantization and event-triggered mechanisms.

This paper presents several new control schemes and a necessary condition for the employment of encryption algorithms for control systems based on event-based communication.

The paper aims to address the combined attitude control and Sun tracking problem in a flexible spacecraft in the presence of external and internal disturbances. The attitude stabilization of a flexible satellite is generally a challenging control problem, because of the facts that satellite kinematic and dynamic equations are inherently nonlinear, the rigid–flexible coupling dynamical effect, as well as the uncertainty that arises from the effect of actuator anomalies.

To deal with these issues in the combined attitude and Sun tracking system, a novel control scheme is proposed based on the adaptive fuzzy Jacobian approach. The augmented spacecraft model is then analyzed and the Lyapunov-based backstepping method is applied to develop a nonlinear three-axis attitude pointing control law and the adaptation law.

Numerical results show the effectiveness of the proposed adaptive control scheme in simultaneously tracking the desired attitude and the Sun.

Reaction wheels are commonly used in many spacecraft systems for the three-axis attitude control by delivering precise torques. If a reaction wheel suffers from an irreversible mechanical breakdown, then it is likely going to interrupt the mission, or even leading to a catastrophic loss. The pitch-axis mounted solar array drive assemblies (SADAs) can be exploited to anticipate such situation to generate a differential torque. As the solar panels are rotated by the SADAs to be orientated relative to the Sun, the pitch-axis wheel control torque demand can be compensated by the differential torque.

The proposed Jacobian control scheme is inspired by the knowledge of Jacobian matrix in the trajectory tracking of robotic manipulators.

The purpose of this paper is to establish the dynamics model of spacecraft during deployment of oblique solar panel using Auto Dynamic Analysis of Mechanical System (ADAMS) and to study the attitude motion of the spacecraft system during the oblique solar panel deployment.

For the case of an oblique solar panel on spacecraft, the dynamics virtual prototype model of deployment of oblique solar panels on spacecraft is established and the dynamics simulation is carried out using ADAMS. The effects of solar panel deployment on the attitude motion of spacecraft with different oblique angles are studied and the attitude motion regularities of spacecraft system are discussed. First, the effects on attitude motion of spacecraft are compared between the normal solar panel deployment and oblique solar panel deployment on a spacecraft. Then the attitude motion of spacecraft during the deployment of solar panel with different oblique angles on spacecraft is studied.

The effects of oblique angle of solar panel deployment on the attitude motion of spacecraft are significant in yaw axis. The bigger the oblique angle, the bigger the changes of yaw angle of spacecraft. However, the bigger the oblique angle, the smaller the changes of roll angle of spacecraft. The effects of oblique angle on pitch angle of spacecraft are slight.

Providing a practical method to study the attitude motion of spacecraft system during deployment of solar panel and improving the engineering application of spacecraft system, which put forward up spacecraft system to the practical engineering.

The paper is a useful reference for engineering design of a spacecraft attitude control system and ground text.

The purpose of this paper is to present an on-orbit frequency identification method for spacecraft directly using attitude maneuver data. Natural frequency of flexible solar arrays plays an important role in attitude control design of spacecraft with solar arrays, and its precision will directly affect the accuracy of attitude maneuver. However, when the flexibility of the solar arrays is large, because of air damping, gravity effect etc., the frequency obtained by ground test shows great error compared with the on-orbit real value. One solution to this problem is to conduct on-orbit identification during which proper identification methods are used to obtain the parameters of interest based on the real on-orbit data of spacecraft.

The observer/Kalman filter identification and eigensystem realization algorithm are used as identification methods, and the attitude maneuver controller is designed using the rigid-body dynamics method.

Two conclusions are drawn in this paper according to results of numerical simulations. The first one is that the attitude controller based on the rigid-body dynamics method is effective in attitude maneuver of the spacecraft. The second one is that the on-orbit parameter identification can be directly achieved by using attitude maneuver data of spacecraft without adding additional missions.

Based on the methods proposed in this paper, it is convenient to obtain the natural frequencies of the spacecraft using the data of the attitude maneuver, which may greatly reduce the cost of on-orbit identification test.

The way of obtaining natural frequencies based on attitude maneuver data of spacecraft provides high originality and value for practical application.

To provide an approach to vibration reduction of flexible spacecraft which operates in the presence of various disturbances, model uncertainty and control input non‐linearities during attitude control for spacecraft designers, which can help them analyze and design the attitude control system.

The new approach integrates the technique of active vibration suppression and the method of variable structure control. The design process is twofold: first design of the active vibration controller by using piezoelectric materials to add damping to the structures in certain critical modes in the inner feedback loop, and then a second feedback loop designed using the variable structure output feedback control (VSOFC) to slew the spacecraft and satisfy the pointing requirements.

Numerical simulations for the flexible spacecraft show that the precise attitude control and vibration suppression can be accomplished using the derived vibration attenuator and attitude control controller.

Studies on how to control the flywheel (motor) under the action of the friction are left for future work.

An effective method is proposed for the spacecraft engineers planning to design attitude control system for actively suppressing the vibration and at the same time quickly and precisely responding to the attitude control command.

This paper fulfills a useful source of theoretical analysis for the attitude control system design and offers practical help for the spacecraft designers.

The purpose of this paper is to present novel robust fault tolerant control design architecture to detect and isolate spacecraft attitude control actuators and reconfigure to redundant backups to improve the practicality of actuator fault detection.

The Robust Fault Tolerant Control is designed for spacecraft Autonomous Rendezvous and Docking (AR&D) using Lyapunov direct approach applied to non‐linear model. An extended Kalman observer is used to accurately estimate the state of the attitude control actuators. Actuators on all three axes (roll/pitch/yaw) sequentially fail one after another and the robust fault tolerant controller acts to reconfigure to redundant backups to stabilize the spacecrafts and complete the required maneuver.

In the simulations, the roll, pitch and yaw dynamics of the spacecraft are considered and the attitude control actuators failures are detected and isolated. Furthermore, by switching to redundant backups, the guarantee of overall stability performance is demonstrated.

A real time actuator failure detection and reconfiguration process using robust fault tolerant control is applied for spacecraft AR&D maneuvers. Finding an appropriate Lyapunov function for the non‐linear dynamics is not easy and always challenging. Failure of actuators on all three axes at the same time is not considered. It is a very useful approach to solve self‐assembly problems in space, spacecraft proximity maneuvers as well as co‐operative control of planetary vehicles in presence of actuator failures.

An approach has been proposed to detect, isolate and reconfigure spacecraft actuator failures occurred in the spacecraft attitude control system. A Robust Fault Tolerant Control scheme has been developed for the nonlinear AR&D maneuver for two spacecrafts. Failures that affect the control performance characteristics are considered and overall performance is guaranteed even in presence of control actuator failures. The architecture is demonstrated through model‐based simulation.

The purpose of this paper is to develop a theoretical design for the alternative attitude control of the rotation about the pitch axis for the nadir-pointing spacecraft in the event of inertial actuator faults.

This paper presents a novel and viable solution to that problem using the combined attitude and sun tracking system (CASTS) that was conceived from an engineering problem-solving toolkit called TRIZ. Linear and fuzzy controllers are used to test the spacecraft CASTS architecture. All the relevant governing equations of the control system and disturbance rejection methods are developed.

The performance of the proposed CASTS control strategy is tested through numerical simulations. The results strongly suggest that the novel proposed control scheme is effective and promising for controlling the satellite attitude and sun tracking simultaneously in the presence of disturbance torques.

This work is mainly focused on the rigid body of the spacecraft hub that contains all attitude control hardware and payload instrumentation, and does not deal with the vibrations evolving from the propellant sloshing and large flexible appendages such as the deployable solar panels and synthetic aperture radar antennas.

The results from this work reveal several practical applications worthy of reducing the weight, size of the spacecraft and, therefore, cost of missions while increasing the instrumentation capabilities.

The proposed CASTS solution is a result of looking much wider than one system from a new combination of attitude control and sun tracking, as well as innovative ways of using it.

The purpose of this paper is to develop a novel nonlinear H_∞ control approach for the nonlinear multivariable attitude tracking of rigid spacecraft.

Based on the transformation of the attitude tracking problem into quaternion error stabilization, the feedback control law is developed by using the normal matrix control theory with the inverse‐additive perturbation description of systems uncertainties, and the Hamilton‐Jacobi‐Isaacs (HJI) partial differential inequality is employed for providing the nonlinear H_∞ control criteria for the proposed control law. The onboard recursive least squares (RLS) estimation algorithm of inertia tensor is used for the further improving of the normal matrix property of the control system. The RLS algorithm is simple enough for the spacecraft borne computer. Computer simulation is performed to demonstrate the effectiveness of the control law proposed.

By the normal matrix control theory, the nonlinear H_∞ control law for attitude tracking is developed without solving the HJI inequality and with the inflight estimation of inertia, the proposed control law is adaptive and robust to the variation of mass properties, and its normality is further improved.

The paper is limited in rigid spacecraft with slowly changing mass property. The flexible influences are not considered.

The paper provides an alternative to the spacecraft researchers/engineers for developing the robust attitude control law with a simple structure and self‐tuning ability.

The paper is the first to provide a robust control based on the normal matrix approach, the HJI inequality, and the estimation of inertia.

The purpose of this paper is to achieve fault reconstruction for reaction wheels in spacecraft attitude control systems (ACSs) subject to space disturbance torques.

Considering the influence of rotating reaction wheels on spacecraft attitude dynamics, a novel non-linear learning observer is suggested to robustly reconstruct the loss of reaction wheel effectiveness faults, and its stability is proven using Lyapunov’s indirect method. Further, an extension of the proposed approach to bias faults reconstruction for reaction wheels in spacecraft ACSs is performed.

The numerical example and simulation demonstrate the effectiveness of the proposed fault-reconstructing methods.

This paper includes implications for the development of reliability and survivability of on-orbit spacecrafts.

This paper proposes a novel non-linear learning observer-based reaction wheels fault reconstruction for spacecraft ACSs.

This paper aims to investigate the problem of finite-time consensus tracking control without unwinding for formation flying spacecraft in the presence of external disturbances.

Two distributed finite-time controllers are developed using the backstepping sliding mode. The first robust controller can compensate for external disturbances with known bounds, and the second one can compensate for external disturbances with unknown bounds.

Because the controllers are designed on the basis of rotation matrix, which represents the set of attitudes both globally and uniquely, the system can overcome the drawback of unwinding, which results in extra fuel consumption. Through introducing a novel virtual angular velocity, exchange of control signals between neighboring spacecraft becomes unnecessary, and it is able to reduce the communication burden.

The two robust controllers can deal with unwinding that may result in fuel consumption by traveling a long distance before returning to a desired attitude when the closed-loop system is close to the desired attitude equilibrium.

Two finite-time controllers without unwinding are proposed for formation flying spacecraft by using backstepping sliding mode. Furthermore, exchange of control signals between neighboring spacecraft is unnecessary.