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1 – 10 of 741The 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…
Abstract
Purpose
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.
Design/methodology/approach
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.
Findings
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.
Practical implications
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.
Originality/value
The paper is a useful reference for engineering design of a spacecraft attitude control system and ground text.
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Abstract
Purpose
The purpose of this paper is to assess the orbital perturbation caused by the gravitational orbit–attitude coupling of spacecraft in the proximity of asteroids.
Design/methodology/approach
The gravitational orbit–attitude coupling perturbation (GOACP), which has been neglected before in the close-proximity orbital dynamics about asteroids, is investigated and compared with other orbital perturbations. The GOACP has its origin in the fact that the gravity acting on a non-spherical extended body is actually different from that acting on a point mass located at the body’s center of mass, which is the approximated model in the orbital dynamics. Besides, a case study of a tethered satellite system is given by numerical simulations.
Findings
It is found that the ratio of GOACP to the asteroid’s non-spherical gravity is the order of ρ/ae, where ρ is the spacecraft’s characteristic dimension and ae is the asteroid’s mean radius. It can also be seen that as ρ increases, GOACP will also increase but the solar radiation pressure (SRP) will decrease due to the decreasing area-to-mass ratio. The GOACP will be more significant than SRP at small orbital radii for a large-sized spacecraft. Based on the results by analyses and simulations, it can be concluded that GOACP needs to be considered in the orbital dynamics for a large-sized spacecraft in the proximity of a small asteroid.
Practical implications
This study is of great importance for the future asteroids missions for scientific explorations and near-Earth objects mitigation.
Originality/value
The GOACP, which has been neglected before, is revealed and studied.
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Pengji Wang and Di Yang
Spacecraft formation flying is a key technology for future astronautics. The relative dynamics of formation flying in eccentric orbits is studied, and the relative motion between…
Abstract
Spacecraft formation flying is a key technology for future astronautics. The relative dynamics of formation flying in eccentric orbits is studied, and the relative motion between spacecrafts is analyzed in this paper. Based on the two‐body problem, the extension of Hill equations is achieved and used in relative dynamics of eccentric orbits. Moreover, the transformation of differential variables is applied, and the algebraic solution of the relative motion is obtained, which can be generally used for spacecraft formation flying in eccentric orbits. In addition, the analysis and numerical simulations are given for the relative motion of spacecraft formation flying. The results demonstrate that each spacecraft in eccentric orbits can run in a periodic motion surrounding the master spacecraft under some conditions. And multiple spacecraft can also set up some special formations according to missions.
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…
Abstract
Purpose
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.
Design/methodology/approach
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.
Findings
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.
Research limitations/implications
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.
Originality/value
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.
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The purpose of this paper is to present a full fourth‐order model of the gravity gradient torque of spacecraft around asteroids by taking into consideration of the inertia…
Abstract
Purpose
The purpose of this paper is to present a full fourth‐order model of the gravity gradient torque of spacecraft around asteroids by taking into consideration of the inertia integrals of the spacecraft up to the fourth order, which is an improvement of the previous fourth‐order model of the gravity gradient torque.
Design/methodology/approach
The fourth‐order gravitational potential of the spacecraft is derived based on Taylor expansion. Then the expression of the gravity gradient torque in terms of gravitational potential derivatives is derived. By using the formulation of the gravitational potential, explicit formulations of the full fourth‐order gravity gradient torque are obtained. Then a numerical simulation is carried out to verify the model.
Findings
It is found that the model is more sound and precise than the previous fourth‐order model due to the consideration of higher‐order inertia integrals of the spacecraft. Numerical simulation results show that the motion of the previous fourth‐order model is quite different from the exact motion, while the full fourth‐order model fits the exact motion very well. The full fourth‐order model is precise enough for high‐precision attitude dynamics and control around asteroids.
Practical implications
This high‐precision model is of importance for the future asteroids missions for scientific explorations and near‐Earth objects (NEOs) mitigation.
Originality/value
In comparison with previous models, a gravity gradient torque model around asteroids that is more sound and precise is established. This model is valuable for high‐precision attitude dynamics and control around asteroids.
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Yuxia Ji, Li Chen, Jun Zhang, Dexin Zhang and Xiaowei Shao
The purpose of this paper is to investigate the pose control of rigid spacecraft subject to dead-zone input, unknown external disturbance and parametric uncertainty in space…
Abstract
Purpose
The purpose of this paper is to investigate the pose control of rigid spacecraft subject to dead-zone input, unknown external disturbance and parametric uncertainty in space maneuvering mission.
Design/methodology/approach
First, a 6-Degree of Freedom (DOF) dynamic model of rigid spacecraft with dead-zone input, unknown external disturbances and parametric uncertainty is derived. Second, a super-twisting-like fixed-time disturbance observer (FTDO) with strong robustness is developed to estimate the lumped disturbances in fixed time. Based on the proposed observer, a non-singular fixed-time terminal sliding-mode (NFTSM) controller with superior performance is proposed.
Findings
Different from the existing sliding-mode controllers, the proposed control scheme can directly avoid the singularity in the controller design and speed up the convergence rate with improved control accuracy. Moreover, no prior knowledge of lumped disturbances’ upper bound and its first derivatives is required. The fixed-time stability of the entire closed-loop system is rigorously proved in the Lyapunov framework. Finally, the effectiveness and superiority of the proposed control scheme are proved by comparison with existing approaches.
Research limitations/implications
The proposed NFTSM controller can merely be applied to a specific type of spacecrafts, as the relevant system states should be measurable.
Practical implications
A NFTSM controller based on a super-twisting-like FTDO can efficiently deal with dead-zone input, unknown external disturbance and parametric uncertainty for spacecraft pose control.
Originality/value
This investigation uses NFTSM control and super-twisting-like FTDO to achieve spacecraft pose control subject to dead-zone input, unknown external disturbance and parametric uncertainty.
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This paper aims to address the problem of formation control for spacecraft formation in elliptic orbits by using local relative measurements.
Abstract
Purpose
This paper aims to address the problem of formation control for spacecraft formation in elliptic orbits by using local relative measurements.
Design/methodology/approach
A decentralized formation control law is proposed to solve the aforementioned problem. The control law for each spacecraft uses only its relative state with respect to the neighboring spacecraft it can sense. These relative states can be acquired by local relative measurements. The formation control problem is converted to n stabilization problems of a single spacecraft by using algebraic graph theories. The resulting relative motion model is described by a linear time-varying system with uncertain parameters. An optimal guaranteed cost control scheme is subsequently used to obtain the desired control performance.
Findings
Numerical simulations show the effectiveness of the proposed formation control law.
Practical implications
The proposed control law can be considered as an alternative to global positioning system-based relative navigation and control system for formation flying missions.
Originality/value
The proposed decentralized formation control architecture needs only local relative measurements. Fuel consumption is considered by using an optimal guaranteed cost control scheme.
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Chengxi Zhang, Hui-Jie Sun, Jin Wu, Zhongyang Fei, Yu Jiang and Guanhua Zhang
This paper aims to study the attitude control problem with mutating orbital rate and actuator fading.
Abstract
Purpose
This paper aims to study the attitude control problem with mutating orbital rate and actuator fading.
Design/methodology/approach
To avoid malicious physical attacks and hide itself, the spacecraft may irregularly switch its orbit altitude within a specific range, which will bring about variations in orbital rate, thereby causing mutations in the attitude dynamics model. The actuator faults will also cause changes in system dynamics. Both factors affect the control performance. First, this paper determines the potential switching orbits. Then under different conditions, design controllers that can accommodate actuator faults according to the statistical law of actuator fading.
Findings
This paper, to the best of the authors’ knowledge, for the first time, introduces the Markovian jump framework to model the possible unexpected mutating of orbital rate and actuator fading of spacecraft and then designs a novel control policy to solve the attitude control problem.
Practical implications
This paper also provides the algorithm design processes in detail. A comparative numerical simulation is given to verify the effectiveness of the proposed algorithm.
Originality/value
This is an early solution for spacecraft attitude control with dynamics model mutations.
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Shima Mousavi and Khashayar Khorasani
A decentralized dynamic neural network (DNN)-based fault detection (FD) system for the reaction wheels of satellites in a formation flying mission is proposed. The paper aims to…
Abstract
Purpose
A decentralized dynamic neural network (DNN)-based fault detection (FD) system for the reaction wheels of satellites in a formation flying mission is proposed. The paper aims to discuss the above issue.
Design/methodology/approach
The highly nonlinear dynamics of each spacecraft in the formation is modeled by using DNNs. The DNNs are trained based on the extended back-propagation algorithm by using the set of input/output data that are collected from the 3-axis of the attitude control subsystem of each satellite. The parameters of the DNNs are adjusted to meet certain performance requirements and minimize the output estimation error.
Findings
The capability of the proposed methodology has been investigated under different faulty scenarios. The proposed approach is a decentralized FD strategy, implying that a fault occurrence in one of the spacecraft in the formation is detected by using both a local fault detector and fault detectors constructed specifically based on the neighboring spacecraft. It is shown that this method has the capability of detecting low severity actuator faults in the formation that could not have been detected by only a local fault detector.
Originality/value
The nonlinear dynamics of the formation flying of spacecraft are represented by multilayer DNNs, in which conventional static neurons are replaced by dynamic neurons. In our proposed methodology, a DNN is utilized in each axis of every satellite that is trained based on the absolute attitude measurements in the formation that may nevertheless be incapable of detecting low severity faults. The DNNs that are utilized for the formation level are trained based on the relative attitude measurements of a spacecraft and its neighboring spacecraft that are then shown to be capable of detecting even low severity faults, thereby demonstrating the advantages and benefits of our proposed solution.
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The purpose of the paper is to present an approach to detect and isolate the sensor failures, using a bank of extended Kalman filters (EKF) using an innovative initialization of…
Abstract
Purpose
The purpose of the paper is to present an approach to detect and isolate the sensor failures, using a bank of extended Kalman filters (EKF) using an innovative initialization of covariance matrix using system dynamics.
Design/methodology/approach
The EKF is developed for nonlinear flight dynamic estimation of a spacecraft and the effects of the sensor failures using a bank of Kalman filters is investigated. The approach is to develop a fast convergence Kalman filter algorithm based on covariance matrix computation for rapid sensor fault detection. The proposed nonlinear filter has been tested and compared with the classical Kalman filter schemes via simulations performed on the model of a space vehicle; this simulation activity has shown the benefits of the novel approach.
Findings
In the simulations, the rotational dynamics of a spacecraft dynamic model are considered, and the sensor failures are detected and isolated.
Research limitations/implications
A novel fast convergence Kalman filter for detection and isolation of faulty sensors applied to the three‐axis spacecraft attitude control problem is examined and an effective approach to isolate the faulty sensor measurements is proposed. Advantages of using innovative initialization of covariance matrix are presented in the paper. The proposed scheme enhances the improvement in estimation accuracy. The proposed method takes advantage of both the fast convergence capability and the robustness of numerical stability. Quaternion‐based initialization of the covariance matrix is not considered in this paper.
Originality/value
A new fast converging Kalman filter for sensor fault detection and isolation by innovative initialization of covariance matrix applied to a nonlinear spacecraft dynamic model is examined and an effective approach to isolate the measurements from failed sensors is proposed. An EKF is developed for the nonlinear dynamic estimation of an orbiting spacecraft. The proposed methodology detects and decides if and where a sensor fault has occurred, isolates the faulty sensor, and outputs the corresponding healthy sensor measurement.
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