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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 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.

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.

The purpose of this paper is to investigate the attitude cooperation control of multi-spacecraft with in-continuous communication.

A decentralized state-irrelevant event-triggered control policy is proposed to reduce control updating frequency and further achieve in-continuous communication by introducing a self-triggered mechanism.

Each spacecraft transmits data independently, without the requirement for the whole system to communicate simultaneously. The local predictions and self-triggered mechanism avoid continuous monitoring of the triggering condition.

This investigation is suitable for small Euler angle conditions.

The control policy based on event-triggered communication can provide potential solutions for saving communication resources.

This investigation uses event- and self-triggered policy to achieve in-communication for the multi-spacecraft system.

This paper aims to investigate the feasibility of using the combination of Lorentz force and aerodynamic force as a propellantless control method for spacecraft formation.

It is assumed that each spacecraft is equipped with several large flat plates, which can rotate to produce aerodynamic force. Lorentz force can be achieved by modulating spacecraft’s electrostatic charge. An adaptive output feedback controller is designed based on a sliding mode observer to account for unknown uncertainties and the absence of relative velocity measurements. Aiming at distributing the control input, an optimal control allocation method is proposed to calculate the electrostatic charge of the Lorentz spacecraft and control commands for the atmospheric-based actuators.

Numerical examples are provided to demonstrate the effectiveness of the proposed control strategy in the presence of J₂ perturbations. Simulation results show that relative motion in a formation can be precisely controlled by the proposed propellantless control method under uncertainties and unavailability of velocity measurements.

The controllability of the system is not theoretically investigated in the current work.

The proposed control method introduced in this paper can be applied for small satellites formation in low Earth orbit.

The main contribution of the paper is the proposal of the propellantless control approach for satellite formation using the combination of Lorentz force and aerodynamic force, which can eliminate the requirement of the propulsion system.

The purpose of this paper is to study and analyse the authorship pattern, degree of collaboration, prepare list of prolific authors and test Lotka’s law of scientific productivity in spacecraft technology research.

Data are collected from the print versions of three journals in the field of spacecraft technology for the period 2001-2011. In all 154 volumes containing 1,907 papers have been analysed, and data are presented in different table headings.

Study reveals that 4,355 authors have contributed 1,907 papers. Journal of Spacecraft and Rockets has published maximum (1,487) number of papers during the study period. Multi-authored papers with 87.15 per cent of contributions have dominated this field of research. Journal of Spacecraft Technology has recorded highest degree of collaboration of 0.90. James M. Longuski has published 20 papers in Journal of Spacecraft and Rockets during the period 2001-2011. Lotka’s law of scientific productivity is tested and conforms only partially.

Study is restricted only for the period 2001-2011, and the data are collected from the print versions of three journals in the field of spacecraft technology research.

As far as space science and technology is concerned, there are not many bibliometric studies reported in the published literature. The present study will add value to the bibliometrics literature and provide publishing trends in spacecraft technology research.

The aim of this paper is to discuss E-sail-based missions (E-sail – electric solar wind sail) towards Venus and Mars. The analysis takes into account the real three-dimensional shape of the starting and arrival orbits and the planetary ephemeris constraints by using the Jet Propulsion Laboratory (JPL) planetary ephemerides model DE405/LE405.

Each mission scenario is parameterized with different values of departure date and spacecraft characteristic acceleration, the latter representing the maximum propulsive acceleration when the Sun–spacecraft distance is 1 au. The transfer trajectories are studied in an optimal framework, using a Gauss pseudospectral method in which the initial guesses for the state and control histories are obtained with a genetic algorithm-based approach.

The paper illustrates the numerical simulations obtained with a spacecraft characteristic acceleration of 1 mm/s², and the results cover a range of launch dates of 17 years for both Earth–Mars and Earth–Venus interplanetary missions. In particular, the numerical results confirm the competitiveness of such a propellantless propulsion system.

A parametric study of the transfer’s flight time corresponding to the optimal departure dates is discussed for different values of the spacecraft characteristic acceleration. The results motivate a further in-depth analysis of the E-sail concept.

This paper extends previous work on optimal trajectories with an E-sail in that the best launch opportunities are investigated. A refined thrust model is also used in all numerical simulations.

In order to succeed in landing asteroids, good accuracy autonomous navigation is absolutely necessary. Aims to describe a new autonomous navigation algorithm.

First, gray images of asteroid surface are acquired by optical navigation camera, and nature feature points are detected and tracked autonomously. Second, the directional vector from spacecraft to the center of each feature point can be computed from the image coordinates in camera focal plane. Then, LIDAR/LRF is directed to three feature points and the distances from spacecraft to feature points are obtained. Last, the relative position vector from spacecraft to the target asteroid is reconstructed base on measurement outputs of navigation cameras and laser light radar (laser range finder).

Suppose the initial conditions presented in this paper, the autonomous optical navigation position error and velocity error are less than 1 m and 0.1 m/s, respectively; this navigation accuracy can satisfy the requirement of soft landing on asteroids.

Based on feature detection and tracking, an autonomous optical navigation scheme is brought out and the validity is confirmed by computer simulation.

This paper aims to study the attitude control problem with mutating orbital rate and actuator fading.

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.

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.

This paper also provides the algorithm design processes in detail. A comparative numerical simulation is given to verify the effectiveness of the proposed algorithm.

This is an early solution for spacecraft attitude control with dynamics model mutations.

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.