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This paper aims to investigate the decentralised strategy to coordinate the reconfiguration of multiple spacecraft.

The system of interest consists of multiple spacecraft with independent subsystem dynamics and local constraints, but is linked through their coupling constraints. The proposed method decomposes the centralised problem into smaller subproblems. It minimises the fuel consumption of multiple spacecraft performing a reconfiguration manoeuvre through an iterative computation. In particular, each agent optimises its individual cost function using the most recently available local solution for the other agents.

The simulation scenarios include spacecraft formation reconfiguration and close manoeuvres around obstacles were conducted. The simulation results showed the fast convergence of the proposed algorithm, while local and inter-vehicle constraints were maintained.

The main advantage of this approach is that it adopts a linear form of the objective function. This allows the local optimisation problem to be formulated as a mixed-integer, linear programming problem, most of which can be quickly solved with resort to commercial software.

The purpose of this paper is to propose strategies for satellite cluster non‐coplanar orbit transfer to reduce fuel cost of formation maintenance and orbit maneuver.

This research tries to use geometric method model to describe the relative motion of satellites in the cluster non‐coplanar orbit transfer, and genetic algorithm (GA) to optimize the proposed maneuver strategies.

Compared with the C‐W equations, the geometric method model is found to be more precise. Three strategies are proposed and optimized to maintain the relative orbit and a strategy of indefinite phase and non‐synchronous costs least fuel.

Geometric method model can be used to describe the relative motion of satellite cluster, especially on elliptical orbits considering the effects of perturbation, with a simple form and good accuracy. Fuel cost minimization is one of the most important issues in formation flight mission.

This paper provides dynamics analysis about formation non‐coplanar orbit transfer, which is involved in minor researches.

The three-axis simulator relies on the air film between the air bearing and the bearing seat to achieve weightlessness and the frictionless motion condition, which is essential for simulating the micro-disturbance torque of a satellite in outer space. However, at the beginning of the experiment, the disturbance torque caused by the misalignment between the center of gravity of the simulator and the center of rotation of the bearing is the most important factor restricting the use of the space three-axis simulator. In order to solve this problem, it is necessary to set the balance adjustment system on the simulator to compensate the disturbance torque caused by the eccentricity. The paper aims to discuss these issues.

In this paper, a study of L1 adaptive automatic balancing control method for micro satellite with motor without other actuators is proposed. L1 adaptive control algorithm adds the low-pass filter to the control law, which in a certain sense to reduce the high-frequency signal and speed up the response time of the controlled system. At the same time, by estimating the adaptive parameter uncertainty in object, the output error of the state predictor and the controlled object can be stabilized under Lyapunov condition, and the robustness of the system is also improved. The automatic balancing method of PID is also studied in this paper.

Through this automatic balancing mechanism, the gravity disturbance torque can be effectively reduced down to 10⁻⁶ Nm, and the automatic balancing time can be controlled within 7 s.

This paper introduces an automatic balancing mechanism. The experimental results show that the mechanism can greatly improve the convergence speed while guaranteeing the control accuracy, and ensuring the feasibility of the large angle maneuver of spacecraft three-axis simulator.

The aim of this paper is to address the problem of fault detection (FD) of linear continuous‐time multi‐agent systems.

A mixed H_∞/H₋ formulation of the FD problem using semi‐decentralized filters is presented.

It is shown that through a decomposition approach the drawbacks of the existing distributed FD design methods in multi‐agent systems can be effectively tackled. An extended linear matrix inequality (LMI) characterization is used to reduce the conservativeness of the design solution by introducing additional matrices in order to eliminate the couplings of the Lyapunov matrices with the agent's matrices.

It is shown that by applying the proposed decomposition approach the FD problem of multi‐agent systems can be solved by analyzing the problem of a set of decoupled systems whose order and complexity are equal to that of a single agent. This procedure will be useful for both simplifying the computational cost of the solution as well as for developing a fault detection filter having a semi‐decentralized architecture.

Application of this methodology to a network of micro‐air vehicles (MAVs) illustrates the effectiveness and capabilities of the proposed design methodology.

The feasibility of the use of reliable and self‐healing network of unmanned systems, cooperative networks, and multi‐agent systems will be significantly enhanced and improved by the development of advanced fault detection and isolation (FDI) technologies.

A semi‐decentralized fault detection (FD) methodology is developed for linear multi‐agent networked systems to reduce the order and complexity of the observers at each agent. A mixed H_∞/H₋ formulation of the FD problem by using semi‐decentralized filters is presented. Using this approach each agent can not only detect its own faults but also is able to detect its nearest neighbor agents’ faults.

Hill equations have definite limitation in the design of multiple spacecraft formation flying in eccentric orbits. To solve the problem, the design method of spacecraft formation flying in a circular reference orbit based on Hill equation can be generalized and applied to spacecraft formation flying in eccentric orbits.

In this paper, T‐H equation is expressed as the explicit function form of reference orbit true anomaly, and the state transition matrix of relative motion of spacecraft formation flying in eccentric orbits is derived. According to the requirement that relative dynamics equation of spacecraft formation flying in eccentric orbits has periodicity solution, the paper theoretically gives the initial condition needed by the long‐term close‐distance spacecraft formation flying including the relationship between relative position and relative velocity. Without perturbation the spacecraft formation, which satisfies the initial periodicity restriction, can keep long‐term close‐distance flying without the need of active control.

Based on the theoretical analysis, some numerical simulations are carried out. The results demonstrate that each spacecraft in eccentric orbits can run in a periodic motion surrounding the center spacecraft under some conditions. And spacecraft formation reconfiguration is implementing according to missions.

Combined with the periodicity restriction primary condition a new method about spacecraft formation reconfiguration is put forward. The method given by this paper can be applied to eccentric orbits of arbitrary eccentricity, and provides theoretical reference for orbit design of spacecraft formation flying in eccentric orbits.

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.

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.

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.

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.

The purpose of this paper is to propose an intuitive and effective cluster flight orbit design method for fractionated spacecraft.

Based on the concept of fractionated spacecraft, orbit design requirements for cluster flight in the case of fractionated spacecraft are proposed, and categorized into three requirements: stabilization requirement, passive safety requirement, and the maximum inter‐satellite distance requirement. These design requirements are then reformulated in terms of relative eccentricity and inclination vectors (E/I vectors) using a relative motion model based on relative orbital elements (ROEs). By using ROEs theory, the cluster flight orbit design issue is modelled as the distribution of relative E/I vectors for each member satellite in the cluster, and solved by combining three different heuristic search methods and one nonlinear programming (NLP) method.

The simulation results show that the NLP method is valid and efficient in solving the cluster flight orbit design problem and that for some cluster flight scenarios, the heuristic search methods can be adopted to give feasible solutions without the NLP method.

The cluster flight scenario in this paper is limited because the cluster should be in the near‐circular low earth orbit (LEO), and the relative distance between the member satellites should be small enough to satisfy the relative motion linearization assumption.

The cluster flight orbit design method proposed in this paper can be applied by fractionated spacecraft mission designers to propose potential cluster flight orbit solutions.

In this paper, the relative E/I vectors method is adopted to propose an intuitive and effective cluster flight orbit design method for fractionated spacecraft.

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 a methodology and experimental results on using global optimization algorithms to determine the optimal orbit, based on the mission requirements, for a set of spacecraft flying in formation with an asteroid.

A behavioral‐based hybrid global optimization approach is used to first characterize the solution space and find families of orbits that are a fixed distance away from the asteroid. The same optimization approach is then used to find the set of Pareto optimal solutions that minimize both the distance from the asteroid and the variation of the Sun‐spacecraft‐asteroid angle. Two sample missions to asteroids, representing constrained single and multi‐objective problems, were selected to test the applicability of using an in‐house hybrid stochastic‐deterministic global optimization algorithm (Evolutionary Programming and Interval Computation (EPIC)) to find optimal orbits for a spacecraft flying in formation with an orbit. The Near Earth Asteroid 99942 Apophis (2004 MN₄) is used as the case study due to a fly‐by of Earth in 2029 leading to two potential impacts in 2036 or 2037. Two black‐box optimization problems that model the orbital dynamics of the spacecraft were developed.

It was found for the two missions under test, that the optimized orbits fall into various distinct families, which can be used to design multi‐spacecraft missions with similar orbital characteristics.

The global optimization software, EPIC, was very effective at finding sets of orbits which met the required mission objectives and constraints for a formation of spacecraft in proximity of an asteroid. The hybridization of the stochastic search with the deterministic domain decomposition can greatly improve the intrinsic stochastic nature of the multi‐agent search process without the excessive computational cost of a full grid search. The stability of the discovered families of formation orbit is subject to the gravity perturbation of the asteroid and to the solar pressure. Their control, therefore, requires further investigation.

This paper contributes to both the field of space mission design for close‐proximity orbits and to the field of global optimization. In particular, suggests a common formulation for single and multi‐objective problems and a robust and effective hybrid search method based on behaviorism. This approach provides an effective way to identify families of optimal formation orbits.

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.