Abstract
The paper describes the preliminary design of a phasing trajectory in a cislunar environment, where the third body perturbation is considered non-negligible. The working framework is the one proposed by the ESA’s Heracles mission in which a passive target station is in a Near Rectilinear Halo Orbit and an active vehicle must reach that orbit to start a rendezvous procedure. In this scenario the authors examine three different ways to design such phasing maneuver under the circular restricted three-body problem hypotheses: Lambert/differential correction, Hohmann/differential correction and optimization. The three approaches are compared in terms of ΔV consumption, accuracy and time of flight. The selected solution is also validated under the more accurate restricted elliptic three-body problem hypothesis.
Highlights
Current and future plans to return to the Moon are considering the presence of a permanent space station in orbit around the L2 Lagrangian point of the Earth–Moon system
The present paper focuses on the phasing trajectory defined within ESA’s Heracles mission, in particular it presents three different methods for the phasing maneuver and compares them
The paper is organized as follows: Section 2 describes the working scenario and relevant CR3BP equations of motion; Section 3 describes three phasing methods and resulting trajectories; a validation using more accurate ER3BP dynamics is presented in Section 4; the discussion and conclusions are described in Sections 5 and 6, respectively
Summary
Current and future plans to return to the Moon are considering the presence of a permanent space station in orbit around the L2 Lagrangian point of the Earth–Moon system. Due to the absence of closed form solution, the numerical propagation of the dynamics requires particular attention to the selection of the boundary conditions To this end, one possible method is based on the solution of the Lambert’s problem. The use of Lambert’s approach could be useful to determine a first guess departing trajectory from the assumed lunar orbit, since the Moon’s gravity can be considered dominant over the Earth’s gravity for altitudes up to 500 km [3] Another approach can be found in [4], where a continuous low thrust maneuver computation is solved via optimal control, with a cost function dependent on the power consumption and time of travel. The paper is organized as follows: Section 2 describes the working scenario and relevant CR3BP equations of motion; Section 3 describes three phasing methods and resulting trajectories; a validation using more accurate ER3BP dynamics is presented in Section 4; the discussion and conclusions are described in Sections 5 and 6, respectively
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