Design of a reconfigurable satellite constellation

This paper provides a fully analytical method to describe a satellite constellation reconfiguration manoeuvre. By making use of low-thrust propulsion and exploiting the Earth’s natural perturbing forces it is possible to analytically describe the reconfiguration of a constellation, achieving a desired separation of both Right Ascension of Ascending Node (RAAN) and Argument of Latitude between satellites. An inherent trade-off exists between the time taken for a manoeuvre and the required ΔV, however the analytical solution presented here allows for a rapid visualisation of the trade-space and determination of the ideal transfer trajectory for a given mission. The general method presented can be applied across a range of scenarios, including constellation deployment and repurposing. The results show that for a scenario with an initial orbit semi-major axis of 6878.14km, and a desired final semi-major axis of 6778.14km it is possible to achieve a separation of 180° argument of latitude between a manoeuvring and a non-manoeuvring reference satellite in approximately 68 hours with a ΔV of 200m/s. To achieve the maximum possible RAAN separation of 90° with a ΔV of 200m/s requires a much longer time of over 218 days. Using two manoeuvring satellites with the same total manoeuvre ΔV was found to be more efficient only for short manoeuvre times. This is quantified and for the case considered it is found that using a 2-satellite manoeuvre is advantageous when changing the argument of latitude and when changing the RAAN