Trajectory corrections for lunar flyby transfers to Sun–Earth/Moon libration point orbits(LPOs)with continuous thrusts are investigated using an ephemeris model.The lunar flyby transfer has special geometrical and dy...Trajectory corrections for lunar flyby transfers to Sun–Earth/Moon libration point orbits(LPOs)with continuous thrusts are investigated using an ephemeris model.The lunar flyby transfer has special geometrical and dynamical structures;therefore,its trajectory correction strategy is considerably different from that of previous studies and should be specifically designed.In this paper,we first propose a control strategy based on the backstepping technique with a dead-band scheme using an ephemeris model.The initial error caused by the launch time error is considered.Since the perturbed transfers significantly diverge from the reference transfers after the spacecraft passes by the Moon,we adopt two sets of control parameters in two portions before and after the lunar flyby,respectively.Subsequently,practical constraints owing to the navigation and propellant systems are introduced in the dynamical model of the trajectory correction.Using a prograde type 2 orbit as an example,numerical simulations show that our control strategy can efficiently address trajectory corrections for lunar flyby transfers with different practical constraints.In addition,we analyze the effects of the navigation intervals and dead-band scheme on trajectory corrections.Finally,trajectory corrections for different lunar flyby transfers are depicted and compared.展开更多
The indirect method for the continuous low-thrust near minimum cumulative longitude orbit transfer problem is addressed.The movement of the satellite is described by the Gauss equation using the modified equinoctial e...The indirect method for the continuous low-thrust near minimum cumulative longitude orbit transfer problem is addressed.The movement of the satellite is described by the Gauss equation using the modified equinoctial elements and replacing time as the system independent variable by the cumulative longitude.The maximum principle is adapted to design the optimal control in order to minimize the final cumulative longitude, and the twopoint-boundary-value problem is derived from the orbit transfer problem.The single shooting method is applied in a numerical experiment, and the simulations demonstrate that the orbit transfer mission is fulfilled and the product of the maximal thrust and the minimum cumulative longitude is near constant.展开更多
基金supported by the Canada Research Chair Program under Grant No.950-230883.
文摘Trajectory corrections for lunar flyby transfers to Sun–Earth/Moon libration point orbits(LPOs)with continuous thrusts are investigated using an ephemeris model.The lunar flyby transfer has special geometrical and dynamical structures;therefore,its trajectory correction strategy is considerably different from that of previous studies and should be specifically designed.In this paper,we first propose a control strategy based on the backstepping technique with a dead-band scheme using an ephemeris model.The initial error caused by the launch time error is considered.Since the perturbed transfers significantly diverge from the reference transfers after the spacecraft passes by the Moon,we adopt two sets of control parameters in two portions before and after the lunar flyby,respectively.Subsequently,practical constraints owing to the navigation and propellant systems are introduced in the dynamical model of the trajectory correction.Using a prograde type 2 orbit as an example,numerical simulations show that our control strategy can efficiently address trajectory corrections for lunar flyby transfers with different practical constraints.In addition,we analyze the effects of the navigation intervals and dead-band scheme on trajectory corrections.Finally,trajectory corrections for different lunar flyby transfers are depicted and compared.
基金supported by the National Natural Science Foundation of China (10832006 60874011)
文摘The indirect method for the continuous low-thrust near minimum cumulative longitude orbit transfer problem is addressed.The movement of the satellite is described by the Gauss equation using the modified equinoctial elements and replacing time as the system independent variable by the cumulative longitude.The maximum principle is adapted to design the optimal control in order to minimize the final cumulative longitude, and the twopoint-boundary-value problem is derived from the orbit transfer problem.The single shooting method is applied in a numerical experiment, and the simulations demonstrate that the orbit transfer mission is fulfilled and the product of the maximal thrust and the minimum cumulative longitude is near constant.
