This paper is concerned with the development and performances of navigation filters for spacecraft relative pose estimation in a mission around an asteroid. For the sake of comparative study, relative pose is represented using two different sets of parameters: 1) conventional Cartesian coordinates along with the attitude quaternion, and 2) the dual quaternion. The spacecraft is equipped with a navigation camera and a laser ranger for position sensing, and with a star tracker and a rate gyroscope for attitude sensing. A highly realistic Truth model is used that includes polyhedron gravity field modelling of small bodies, polyhedron gravity gradient torque modeling, advanced hardware modeling of the navigation camera and laser ranging observations and errors. Results of extensive Monte-Carlo simulations for various orbital scenarios around two different asteroids, Kleopatra and Itokawa, will be shown. It turns out that not only are the filters able to estimate the relative pose with high accuracy, but that the gyroscope drift and asteroid angular rates have good observability too. The latter depends on the frequency and geometry of the landmarks lines-of-sight detected within the camera field-of-view. In some cases, the dual quaternion filter shows a better transient, when compared to the conventional filter. Their steady-state accuracies are in general similar. The trade-offs in using dual quaternion filters as opposed to standard filters are discussed.

A method is developed to determine station-keeping maneuvers for a fleet of satellites collocated in a geostationary slot. The method is enabled by a linear time-varying formulation of the satellite orbit dynamics in terms of nonsingular orbital elements. A leader–follower control hierarchy is used, such that the motion of the follower satellites is controlled relative to the leader. Key objectives of the station-keeping method are to minimize propellant consumption and to limit the number of maneuvers while guaranteeing safe separation between the satellites. The method is applied in a realistic simulation scenario, including orbit determination, as well as actuation and modeling errors. The method is demonstrated to work for a fleet of four satellites with differences in mass, surface area, and propulsion system for a maneuver cycle of one week. It is then demonstrated that, by reducing the maneuver cycle duration to one day, the method allows collocation of 16 satellites in a single slot, without penalties on propellant consumption.