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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.

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.

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.

An unexpected space debris population was recently identified as being very high area-to-mass ratio (HAMR)\ud debris in geostationary orbit (GEO). Scientists hypothesised about their origins, consensus is that these objects are\ud multi-layer insulation (MLI), separated from the spacecraft due to fragmentation events or delamination. These\ud objects, like any other debris, pose a collision hazard for active satellites. Debris are usually considered rigid bodies;\ud this assumption, however, is not fully justified for the MLI, as it lacks almost any structural strength. Large sheets of\ud MLI are very flexible, and the change in geometry due to their flexibility can affect the effective area-to-mass ratio\ud (AMR). This will cause varying effects due to external or internal forces and moments such as solar radiation\ud pressure SRP, atmospheric drag, electromagnetic fields, or centrifugal and Coriolis forces. This, in turn, will affect\ud the evolution of the orbital parameters, in a way that is difficult to predict. This paper introduces a simplified but\ud effective model to represent the deformation of such debris, subject in particular to torques caused by solar radiation\ud pressure and the Earth gravitational field, by means of Finite Element Method (FEM). This model adds a further set\ud of dynamical equations, which accounts for the flexibility of the object, into the attitude and orbital equations; the\ud resulting system is then numerically integrated to better evaluate the coupling between orbital and attitude dynamics.\ud Due to a more precise estimation and prediction of the actual shape and orientation of the debris at any given time,\ud than by simply assuming the case of a rigid body, the effects of the perturbations on the orbit can be computed more\ud precisely, leading to improvements for the long-term prediction, over 150 days, of the orbital evolution. Results\ud show that for debris in GEO the eccentricity change for flexible debris is different than for equivalent rigid bodies,\ud and their attitude motions are unique.