Distributed Spacecraft Path Planning and Collision Avoidance via Reciprocal Velocity Obstacle Approach

This paper presents the development of a combined linear quadratic regulation and reciprocal velocity obstacle (LQR/RVO) control algorithm for multiple satellites during close proximity operations. The linear quadratic regulator (LQR) control effort drives the spacecraft towards their target position while the reciprocal velocity obstacle (RVO) provides collision avoidance capabilities. Each spacecraft maneuvers independently, without explicit communication or knowledge in term of collision avoidance decision making of the other spacecraft in the formation. To assess the performance of this novel controller different test cases are implemented. Numerical results show that this method guarantees safe and collision-free maneuvers for all the satellites in the formation and the control performance is presented in term of Δv and fuel consumption

Modified Nonlinear Integral Sliding Mode Control for Satellite Attitude Stabilization Using Magnetically Suspended Gimbaled Momentum Wheel

This paper treats the attitude stabilization problem for satellite using only one MSGMW (Magnetically Suspended Gimbaled Momentum Wheel). To start, the coupled dynamic model of satellite and MSGMW is defined and simplified based on the fact that the attitude errors are small during the mission mode that the MSGMW services. In order to improve the dynamic performance, reduce the steady state error and avoid the chattering phenomenon, a modified integral chattering-free sliding mode controller with a nonlinear integral function and a saturation function is introduced. Lyapunov theory is employed to prove the convergence characteristic outside the boundary layer and the terminal convergence characteristic inside the boundary layer. A numerical simulation example is employed to show the effectiveness and suitability of the proposed controller

Interception and deviation of near Earth objects via solar collector strategy

A solution to the asteroid deviation problem via a low-thrust strategy is proposed. This formulation makes use of the proximal motion equations and a semi-analytical solution of the Gauss planetary equations. The average of the variation of the orbital elements is computed, together with an approximate expression of their periodic evolution. The interception and the deflection phase are optimised together through a global search. The low-thrust transfer is preliminary designed with a shape based method; subsequently the solutions are locally refined through the Differential Dynamic Programming approach. A set of optimal solutions are presented for a deflection mission to Apophis, together with a representative trajectory to Apophis including the Earth escape

Study on stability and rotating speed stable region of magnetically suspended rigid rotors using extended Nyquist criterion and gain-stable region theory

This paper presents a novel and simple method to analyze the absolute stability and the rotor speed stable region of a magnetically suspended rotor (MSR). At the beginning of the paper, a complex variable is introduced to describe the movement of the MSR and a complex coefficient transfer function is obtained accordingly. The equivalent stability relationship between this new variable and the two traditional deflection angles is also demonstrated in a simple way. The detailed characteristics of the open-loop MSR system with time delay are studied carefully based on the characteristics of its Nyquist curve. A sufficient and necessary condition of absolute stability is then deduced by using an extended complex Nyquist stability criterion for MSRs. Based on the relationship between the rotor speed and gain-stable region proposed in this paper, the rotor speed stable region can be solved simply and directly. The usefulness and effectiveness of the proposed approaches are validated by examples and simulations

Electromagnetic Formation Flying with Eccentric Reference Orbits

Over the last decade, a considerable amount of research work has been done in the area of spacecraft formation flight, with particular emphasis on control techniques using thruster-based systems. Nevertheless, thrusters require propellant to work and this limit the lifetime of the mission. Electromagnetic Formation Flight (EMFF) is presented in this paper as a fuel-less strategy to control spacecraft formations by means of electromagnets. In EMFF, spacecraft can be equipped with one or more coils and reactions wheels which could be arranged in several combinations according to mission requirements. An electric current flows through the coils in order to produce a magnetic dipole in a specific direction. The magnetic field of a spacecraft reacts against the magnetic dipoles of the others, generating forces and torques which in turn could be used as control inputs. The main objective of this paper is to provide a formulation for EMFF when a formation is moving in eccentric reference orbits and for this purpose, the Tschauner and Hempel model will be used. Results are presented after analysing different formation scenarios providing the necessary magnetic requirements for station keeping and resolving which cases are suitable to be controlled by this technology. High-Temperature Semiconductor (HTS) plays an important role in EMFF and for that reason the paper also investigates the correlation of the magnetic force and the coil mass, which in turn affects the total mass of the spacecraft

Autonomous pointing control of a large satellite antenna subject to parametric uncertainty

With the development of satellite mobile communications, large antennas are now widely used. The precise pointing of the antenna’s optical axis is essential for many space missions. This paper addresses the challenging problem of high-precision autonomous pointing control of a large satellite antenna. The pointing dynamics are firstly proposed. The proportional–derivative feedback and structural filter to perform pointing maneuvers and suppress antenna vibrations are then presented. An adaptive controller to estimate actual system frequencies in the presence of modal parameters uncertainty is proposed. In order to reduce periodic errors, the modified controllers, which include the proposed adaptive controller and an active disturbance rejection filter, are then developed. The system stability and robustness are analyzed and discussed in the frequency domain. Numerical results are finally provided, and the results have demonstrated that the proposed controllers have good autonomy and robustness

Orbital dynamics of large solar power satellites

Designs for geostationary SPS are extremely large in scale, more than an order of magnitude larger than the International Space Station. The problem of how to control the orbital motion of such large structures, accounting for various perturbing forces, is therefore a topic worthy of further study. The primary objective of the proposed research is to perform a detailed study of SPS orbit dynamics, obtaining a comprehensive understanding of the effect of perturbations on orbits of large SPS structures over a time-frame commensurate with proposed SPS lifetimes (30-40 years). Analytical equations derived by the process of averaging of the SPS equations of motion shall be used in determining the long-term orbital behaviour. Previous studies have simply assumed a geostationary orbit (GEO) then designed control systems for maintaining it thus. It is found that an alternative SPS orbital location known as the geosynchronous Laplace plane (GLP) is superior to GEO. An SPS in GLP requires virtually no fuel to maintain its orbit, avoids the main orbital debris population originating from GEO satellites and is extremely robust, i.e. loss of control is inconsequential. The GLP SPS saves of order 10<sup>4</sup> to 10<sup>5</sup> kg per year in fuel compared to a GEO SPS for equivalent power delivery compared to GEO

Towards Designing a Credible Hazardous NEA Mitigation Campaign of Dual-deflection Act

Given a limited warning time, an asteroid impact mitigation campaign would hinge on uncertainty-based information consisting of remote observational data of the identified Earth-threatening object, general knowledge on near-Earth asteroids, and engineering judgment. Due to these ambiguities, the campaign credibility could be profoundly compromised. It is therefore imperative to comprehensively evaluate the inherent uncertainty in deflection and plan the campaign accordingly to ensure successful mitigation. This research demonstrates dual-deflection mitigation campaigns consisting of primary and secondary deflection missions, where both deflection performance and campaign credibility are taken into consideration. The results of the dual-deflection campaigns show that there are trade-offs between the competing aspects: the total interceptor mass, interception time, deflection distance, and the confidence in deflection. The design approach is found to be useful for multi-deflection campaign planning, allowing us to select the best possible combination of deflection missions from a catalogue of various mitigation campaign options, without compromising the campaign credibility

Hazardous Near Earth Asteroid Mitigation Campaign Planning Based on Uncertain Information on Asteroid Physical Properties

Given a limited warning time, an asteroid impact mitigation campaign would hinge on uncertainty-based information consisting of remote observational data of the identified Earth-threatening object, general knowledge on near-Earth asteroids (NEAs), and engineering judgment. Due to these ambiguities, the campaign credibility could be profoundly compromised. It is therefore imperative to comprehensively evaluate the inherent uncertainty in deflection and plan the campaign accordingly to ensure successful mitigation. This research demonstrates dual-deflection mitigation campaigns consisting of primary and secondary deflection missions, where both deflection performance and campaign credibility are taken into consideration. The results of the dual-deflection campaigns show that there are trade-offs between the competing aspects: the total interceptor mass, interception time, deflection distance, and the confidence in deflection. The design approach is found to be useful for multi-deflection campaign planning, allowing us to select the best possible combination of deflection missions from a catalogue of various mitigation campaign options, without compromising the campaign credibility

A comparative assessment of different deviation strategies for dangerous NEO

In this paper a number of deviation strategies for dangerous Near Earth Objects (NEO) have been compared. For each strategy (i.e. Solar Collector, Nuclear Blast, Kinetic Impactor, Low-thrust Propulsion, Mass Driver) a multi criteria optimisation method has been used to reconstruct the set of Pareto optimal solutions minimising the mass of the spacecraft and the warning time, and maximising the deviation. Then, a dominance criterion has been defined and used to compare all the Pareto sets. The achievable deviation at the MOID, either for a low-thrust or for an impulsive variation of the orbit of the NEO, has been computed through a set of analytical formulas. The variation of the orbit of the NEO has been estimated through a deviation action model that takes into account the wet mass of the spacecraft at the Earth. Finally the technology readiness level of each strategy has been used to compute a more realistic value for the required warning time