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Based on the analytical solutions of T-H equations and its state transition matrix form, the open-loop control method of spacecraft impulsive relative hovering was studied， which is promising for practical engineering use. The true anomaly intervals of the hovering impulse were optimized by the nonlinear mathematical programming. Based on the calculation of collision probability, the method of safety analysis and risk management was proposed. The numerical simulations show that the introduced relative hovering method can be used for circular and elliptical reference orbits hovering. Furthermore, the local optimal solution can be obtained by applying the true anomaly intervals optimization method. The maximum collision probability and the minimum relative distance nearly appear at the same time. And, the smaller the relative distance is, the larger the collision probability. [ABSTRACT FROM AUTHOR]

This paper provides an optimal entry trajectory of the winged Mars probe, right after it changed the orbit into the Martian atmosphere braked by reverse thrust engine. This paper establishes Martian atmosphere model and gravitational field model corresponding on the density of the Martian atmosphere and gravitational field parameters, and determined the dynamic equations of the Mars probe. Within all the constraint conditions in the entering process, this paper designs an optimal entry trajectory based on the genetic algorithm. In addition, this paper provides different thrust engine scheme through comparing them. The results show that the second scheme can achieve the optimal performance index and the landing speed is 13.6 m/s, ultimately realizing the minimum consumption of energy and aerodynamic heat emission structure quality in the process of entering and soft landing. [ABSTRACT FROM AUTHOR]

A new concept of dynamic reentry footprint was presented to ensure the reentry of the spacecraft safely and accurately. Concerned with all the given flight states and the constraint conditions of the spacecraft during the whole reentry flight process, the maximum reachable area on the earth surface can be calculated by the given numerical algorithm. Combined with the dynamic reentry numerical algorithm, the modified genetic algorithm was applied to optimize the reentry flight trajectories, and all the qualified trajectories which constitute the landing footprint can be calculated. Through a series of reentry flight numerical simulations, the characteristics of dynamic reentry footprint was summarized, which can be used to optimize the guidance and flight trajectory. [ABSTRACT FROM AUTHOR]

To satisfy the requirement of long-term on-orbit displaced spacecraft for specific tasks, a hybrid propulsion technology, consisting of solar electric propulsion (SEP) and solar sail, was studied for the displaced geostationary orbit. The relation between the sail attitude angles and the SEP acceleration for hybrid propulsion system was firstly deduced. A single variable extremum problem obtained from the first-order optimality condition was solved to minimize the SEP acceleration and acquire the control law of solar sail attitude angles. Numerical simulations show that the hybrid propulsion approach can save about 57% of fuel consumption per year compared with the SEP method. The hybrid propulsion approach achieves the goal of improving the lifetime of spacecraft and meets the requirements of the displaced geostationary orbits. [ABSTRACT FROM AUTHOR]

A general non-smooth trajectory optimization approach was presented. The mesh refinement procedure based on data compress on dyadic grids was improved to increase the efficiency, usability and flexibility, and was then combined with the local collocation method which took advantage of scaling, sparsity and different numerical differentiation of nonlinear programming (NLP). Numerical simulation on Earth-to-Mars transfer orbit optimization demonstrates that the approach captures non-smooth regions exactly with density mesh and has the potential of on-board optimization. Furthermore, the approach captures non-smooth regions of states even though the mesh is refined according to the control variables. For Earth-to-Mars transfer orbit optimization, it is better to define the thrust angle in [-0, 360°) than that in [-180°, 180°). [ABSTRACT FROM AUTHOR]

The performances of the ADAM-type mast, such as mass, packaging size, stiffness and strength, are determined by its configuration parameters. A new design of configuration parameters based on maximizing specific stiffness was presented. By MATLAB and ANSYS joint simulation, the varying of specific stiffness with different configuration parameters was analyzed. The simulation results show that the optimum value of the length ratio of longeron to batten and the canister radius increases with the increase of the length of mast, when the length of ADAM-type varies in the range of 20 to 100 meters. While the length of mast is 20, 60, and 100 meters, the optimum value of the canister radius varies from 0.3 to 0.4 meter, 0.5 to 0.7 meter and 0.6 to 0.9 meter respectively. The design flow of configuration parameter based on maximizing specific stiffness is as follow: the canister radius is settled at first, then the length ratio of longeron to batten should be determined, and the number of longitudinal member could he finally settled. By this method, configuration parameters of the mast in the SRTM program are designed, and the results approximate to that of ADAM's. [ABSTRACT FROM AUTHOR]

To realize the trajectory optimization of the lunar module in ascent stage, the dynamical model of the lunar module was established and nondimensionalized to construct the optimal control-oriented model. Based on Pontryagin's minimal principle and taking the fuel consumption as the optimizing index, the optimal problem was converted into a time free two>point boundary value problem(TPBVP). Using one of initial value hypothesizing methods, combined with the forward sweep algorithm, the TPBVP was solved and the optimized trajectory was obtained. Numerical simulation was conducted. The simulation results justify the rapid convergence rate and high reliability of the proposed method. [ABSTRACT FROM AUTHOR]

At present, most of the research on low thrust escape trajectory optimization were limited to planar escape or unconstrained engine model. Using the indirect method, the unknown parameters of the two-point-boundary-value-problem were bounded in a hyper unit sphere by multiplying a positive unknown parameter to the fuel optimal performance index. Firstly, the two-dimensional short-time fuel-optimal escape trajectory without control constraint was solved by the homotopy method and curve fits technology. Finally, the three-dimensional long-time fuel-optimal escape trajectory with control constraint was solved. This method is validated to be high-accuracy, good-convergence and efficient for low thrust escape trajectory optimization. [ABSTRACT FROM AUTHOR]

The Earth-Mars low-thrust trajectory was optimized by Artificial Immune Algorithm. Firstly, the system model of Earth-Mars low-thrust trajectory was established and normalized. Secondly, the optimization problem of low-thrust trajectory was converted into a nonlinear programming via parameterization and penalty function. Finally, a GAIA (Guiding Artificial Immune Algorithm) was proposed to solve this optimization problem. Simulation results show that the GAIA has fast convergence rate and high optimization precision. Moreover, GAIA algorithm avoids many shortcomings, such as initial value sensitivity, ill-conditioned gradient and local convergence. The effectiveness of GAIA for low-thrust trajectory optimization is validated. [ABSTRACT FROM AUTHOR]