Application of low thrust propulsion to interconnect ballistic trajectories on invariant manifolds associated with multiple circular restricted three body systems has been investigated. Sun-planet three body models have been coupled to compute the two ballistic trajectories, where electric propulsion is used to interconnect these trajectories as no direct intersection in the Poincarè sections exists. The ability of a low thrust to provide the energy change required to transit the spacecraft between two systems has been assessed for some Earth to Mars transfers. The approach followed consists in a planetary escape on the unstable manifold starting from a periodic orbit around one of the two collinear libration points near the secondary body. Following the planetary escape and the subsequent coasting phase, the electric thruster is activated and executes an ad-hoc thrusting phase. The complete transfer design, composed of the three discussed phases, and possible applications to Earth-Mars missions is developed where the results are outlined in this paper.
Of the three collinear libration points of the Sun-Earth Circular Restricted Three-Body Problem (CR3BP), L 3 is that located opposite to the Earth with respect to the Sun and approximately at the same heliocentric distance. Whereas several space missions have been launched to the other two collinear equilibrium points, i.e. L 1 and L 2 , taking advantage of their dynamical and geometrical characteristics, the region around L 3 is so far unexploited. This is essentially due to the severe communication limitations caused by the distant and permanent opposition to the Earth, and by the gravitational perturbations mainly induced by Jupiter and the close passages of Venus, whose effects are more important than those due to the Earth. However, the adoption of a suitable periodic orbit around L 3 to ensure the necessary communication links with the Earth, or the connection with one or more relay satellites located at L 4 or L 5 , and the simultaneous design of an appropriate station keeping-strategy, would make it possible to perform valuable fundamental physics and astrophysics investigations from this location. Such an opportunity leads to the need of studying the ways to transfer a spacecraft (s/c) from the Earth's vicinity to L 3 . In this contribution, we investigate several trajectory design methods to accomplish such a transfer, i.e., various types of two-burn impulsive trajectories in a Sun-s/c two-body model, a patched conics strategy exploiting the gravity assist of the nearby planets, an approach based on traveling on invariant manifolds of periodic orbits in the Sun-Earth CR3BP, and finally a low-thrust transfer. We examine advantages and drawbacks, and we estimate the propellant budget and time of flight (TOF) requirements of each.
By replacing the liquid metal propellant with a ionic liquid, it is possible to develop a new, simplified FEEP system that combines most of the heritage and the advantages of the linear slit geometry with the easy of handling and operation of a more benign propellant. In view of the development of such Ionic Liquid FEEP thruster (IL-FEEP), an internal development activity is underway at Alta, aimed at the design and testing of an innovative linear slit thruster derived from the cesium experience. This paper presents the results of recent experimental campaigns aimed at assessing the performance of linear slit FEEP emitters fed with a ionic liquid propellant. For the first time, beam composition was evaluated using a time-of-flight mass spectrometry technique, allowing for a reliable estimate of the thruster’s specific impulse
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