The PLEIADES mission, an Earth observation mission, is composed of two satellites orbiting on the same Low Earth orbit. The reference orbit is sun-synchronous, phased, and has a frozen eccentricity. The two satellites are separated by 180° in argument of latitude. At the end of the mission, mission analysis advocates to deorbit the satellites one by one, while following a two-part deorbiting strategy:

• First, manoeuvres are commanded for deorbiting PLEIADES satellite while maintaining the eccentricity constant during the descent
• Then, when a given altitude is reached, the perigee is lowered to make the satellite burn into the atmosphere.

In case of a battery failure causing a quick power leakage, the satellite under consideration must be deorbited as quickly as possible while trying to limit battery discharge. The deorbiting must also be carried out quickly to free the operational orbit and comply as much as possible with the criteria defined by the French Space Operations Act. Consequently, solar panels must be pointed toward the Sun when the satellite is on the illuminated portion of the orbit. This pointing strategy is not compatible with the braking manoeuvres. Hence, the manoeuvres needed for deorbiting can only be commanded during the eclipse phase. This means that a perfect circular descent cannot be carried out, and eccentricity will no longer be constant, as the eclipse phase is shorter than half an orbit. This raises collision risks between the two PLEIADES satellites, as the period of the deorbiting satellite will change, and its argument of apogee will start to drift toward the argument of perigee of the other satellite due to the degraded eccentricity.
 
In this paper, we will demonstrate that there is an optimal deorbiting strategy that allows deorbiting one PLEIADES satellite in minimal time (under 24h), while both avoiding collision risk with the other satellite and minimizing battery discharge.

We will first demonstrate that the deorbiting criterion to meet is maintaining the apogee of the deorbited PLEIADES satellite lower than the reference perigee. We will then prove that this criterion is always met after a certain number of braking manoeuvres. We will also determine analytically the minimum number of manoeuvres, by considering first a Keplerian motion and assuming that propulsion parameters are constant between manoeuvres. In particular, we will show that this minimum value is reached when finding the right equilibrium between the magnitude of each manoeuvre and the associated eccentricity degradation. We will finally test the robustness of the solution against drag effect, AOCS actuators constraints and propulsion parameters update between each manoeuvre. This analysis will be carried out by using a semi analytical propagation based on STELA (Semi-analytic Tool for End of Life Analysis) software.