Satellite deorbiting by means of electrodynamic tethers Part II: System

This paper aims to assess the efficiency of a de-orbiting system based upon conductive tethers under realistic assumptions for its interaction with the ionospheric environment. We analyze the configuration made up of a 2.5 - 10 km tether, a passive inflatable collector of 2.5 - 10 m radius at the positive termination and a hollow cathode at the negative one. Voltages and current in the system are computed from the equation of the equivalent circuit, making use of the IRI-90 ionospheric model. The resulting electromagnetic drag forces have been used to compute the evolution of the orbital elements (especially the semi-major axis) and the re-entry times. Our results indicate that a typical satellite of 500 kg mass at 1300 km altitude can de-orbit in 20-100 days, for a broad range of orbital inclinations and solar activity. The validity of the concept is further strengthened by the comparison with alternative propulsion systems.