Trajectory Design And Optimization For Cycler Architecture

A cycler trajectory is a repeatable round-trip trajectory that shuttles between any two or more celestial bodies. The cycler architecture provides a reusable vehicle that once placed in orbit can shuttle crews and cargo between planets using very little fuel, which is regarded as an alternative approach to developing and maintaining large-scale space transportation with long duration and low cost. Taking trajectories for cycler architecture as investigative object, this dissertation studies the cycler trajectory design, transfer trajectory design, parking orbit selection, hyperbolic rendezvous. The main results achieved in this dissertation are summarized as follows.Cycler trajectories in the Earth-Moon system are designed. According to arcs that belong to the generating orbit of the second species and direction of motion, five types of cycler trajectories are designed by patched conic technique based on planar circular restricted three body model. Characteristics of five types of cycler trajectories are analyzed from the aspects of period time, perigee distance, perilune distance, the speed of rendezvous and docking, stability of the trajectories. Proper type of trajectories are selected for the mission considering the engineering constraints of cycler architecture. Then the trajectory selected from the proper type is optimized to meet the requirements of the mission.Trajectory corrections for Earth-Mars cyclers are studied. In the simplified model, the gravities of the Earth and Mars are introduced successively, then trajectory corrections are implemented step by step. the two strategies of maneuvers near the aphelion and near the planets are compared. Numerical results on inbound Aldrin cycler are obtained for a 15-year period from 2031 through 2045.The optimum two-impulse transfer problem between a low Earth orbit and sample cycler orbits in the framework of the circular restricted three-body problem is explored. Optimum two-impulse orbital transfers between inclined elliptical orbits is solved by numerical methods and the results are served as initial guesses for determining optimal impulsive escape trajectories in the circular restricted three-body problem. Finally the transfer trajectories are optimized by shooting methods and non-linear programming in the circular restricted three-body problem.The injection into the Earth-Mars cycler is divided into three phases: spiraling out, planetary flybys, and spinning. The technology of V? leveraging is applied to establish cycler vehicles in their desired orbits.A method to select the spacecraft’s orbit around a planet for a roundtrip mission is described. Taking the natural precession of the parking orbit during the stay time into consideration, the solution existence of parking orbits for five maneuver strategies is discussed. The five maneuver strategies are comparatively analyzed in a practical mission.The tangent technology is applied to the two-impulse hyperbolic rendezvous for improved rendezvous safety. For a given rendezvous point, the solution to the optimal elliptic transfer orbit with a terminal tangent burn between circular and hyperbolic orbits is provided in a closed-form expression. By calculating all the optimal elliptic transfer orbits with feasible rendezvous points, a transfer corridor is found in which the total velocity increments are significantly less than that outside. In addition, the impact of time constraints including docking time and transfer time on the window for rendezvous mission is analyzed. According to the existence of the window for rendezvous mission, two separate two-impulse strategies are proposed, one is tangent rendezvous and tangent separation, the other is tangent rendezvous and non-tangent separation. Finally, the optimization results of several examples demonstrate the feasibility of the two strategies.