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To satisfy the requirement of contingency rescue in the lunar orbit insertion phase for crewed lunar exploration missions, this paper designs the contingency point return trajectory. First, according to different failure situations of the main propulsion system in the lunar orbit insertion phase, different types of contingency return modes are analyzed and summarized. Second, considering the terminal constraints of point return, a calculation model of the contingency point return trajectory is established, and a serial optimization process is adopted to obtain a high-fidelity solution. Finally, a numerical simulation is used to verify the model. The simulation results indicate that this model can be applied to determine the contingency point return trajectories under different situations in the lunar orbit insertion phase and demonstrate the validity and feasibility of the trajectory calculation model proposed in this paper. The research conclusions can provide a reference for the design of contingency rescue schemes in the lunar orbit insertion phase for crewed lunar exploration missions. [ABSTRACT FROM AUTHOR]

A direct abort orbit design method is presented for direct abort missions in the Earth-Moon transfer phase of crewed lunar exploration missions. First, according to the demand of an emergency rescue in the Earth-Moon transfer phase, two direct abort orbit schemes are introduced. Then, a serial orbit design method is proposed for a high-fidelity direct abort orbit. An analytical model is established for the calculation of initial values, and the optimization design is performed in the high-fidelity orbit model to determine a single-impulse abort orbit. A hybrid optimization design process is proposed to generate a two-impulse abort orbit. The results of simulation examples verify the validity and feasibility of the proposed direct abort orbit design method. Finally, extensive simulations are carried out to analyze the characteristics of abort impulse and abort return time and reveal the general rules of direct abort orbits. The research conclusions can provide a reference for the design of emergency rescue schemes in future crewed lunar exploration missions. [ABSTRACT FROM AUTHOR]

A three-impulse point return orbit design method is proposed for the manned lunar high-latitude exploration mission. First, considering the restrictions of the reentry corridor and the landing point simultaneously, the constraint conditions of the point return orbit are analyzed. Second, a serial orbit design strategy is presented. In the initial calculation, by decoupling the three-impulse transfer and the lunar escape, a two-segment patched method based on the pseudo-perilune parameters is proposed to determine the three-impulse point return orbit. The three-impulse transfer segment is backward designed by combining maneuvering at the special points with the Lambert problem in the hybrid orbit model, while the lunar escape segment is forward solved by introducing a comprehensive conic method. In the accurate solution, the result of the initial calculation serves as the initial value to perform the backward and forward integration for two segments in the high-fidelity model, respectively. The simulation examples verify the effectiveness and feasibility of this strategy, and show the advantages of high accuracy of the initial calculation and low velocity increment consumption. Finally, a large amount of simulations are conducted by applying this strategy, and the orbit characteristics are discussed. The research conclusions can provide a reference for the design of the point return orbit scheme for the manned lunar exploration mission. • Two-segment patched method for designing the three-impulse point return orbit. • New three-impulse transfer method with low impulse consumption. • Comprehensive conic method for generating the lunar escape segment. [ABSTRACT FROM AUTHOR]

A near rectilinear halo orbit (NRHO) is regarded as a potential orbit for a future deep space station that can effectively support sustainable crewed lunar exploration missions. In this paper, the L 2 southern NRHOs are selected as the research object, and a direct transfer trajectory from an NRHO to a low lunar orbit (LLO) is designed and analyzed. First, based on the circular restricted three-body problem, the characteristics of NRHOs are discussed including geometric behaviour, stability and synodic resonance. Second, optimal direct transfer trajectories are obtained by combining a local gradient optimization with a numerical continuation strategy. A reachable domain calculation model is established. Finally, simulations are carried out to verify the feasibility of the trajectory design method. The relationship between the velocity change and the reachable domain is further analysed through simulation calculations. [ABSTRACT FROM AUTHOR]

Lunar exploration architecture can be made more flexible and reliable with the support of a low-Earth orbit (LEO) space station. This study therefore evaluated a proposed hybrid optimization scheme to design the entire trajectory of a reusable spacecraft starting from trans-Earth injection (EI) at the perilune and ending at an LEO space station. As such a trajectory has multiple constraints and multiple dynamical models, it is divided into the trans-Earth phase, aerocapture phase, and postatmospheric phase. The optimization scheme is performed at two levels: sublevel and top level. At the sublevel, two novel pseudo rules are proposed to optimize the trans-Earth trajectory so that it satisfies the coplanar constraints of the space station. Then, in the aerocapture phase, the bank angle is optimized to satisfy the mission constraints, and in the atmospheric phase, the one-impulsive maneuver is performed and optimized to insert the spacecraft into the target space station orbit. The multiple phases are connected to each other by boundary conditions where the terminal state of the previous phase is transformed into the initial state of the following phase. At the top level, the vacuum perigee height is selected as the mission design variable based on problem characteristics analysis and a hybrid optimization scheme is conducted to minimize the total velocity increment. The simulation results demonstrate that the proposed hybrid optimization method is effective for the design of an entire trajectory with acceptable velocity cost which is less than that in the previous study. The coplanar constraints of the space station and other mission constraints in each phase are also satisfied. Furthermore, the proposed trajectory design method is shown to be applicable to a reusable spacecraft returning to an LEO space station parked in any arbitrary orbital plane. [ABSTRACT FROM AUTHOR]