Research On Optimization Technology Of Gravity-Assist Trajectory With Maneuvers In Deep Space Exploration

Deep space exploration is significant for China's aerospace development.In planning of actual deep space missions,a set of fast and efficient trajectory optimization method is of great importance.For deep space missions further than the Mars,launch energy of existing launch vehicle is difficult to complete direct interplanetary transfer of sufficient mass payload.As a result,celestial gravity assist technology has been widely used in orbit design.This dessertation mainly study on sequence searching algrithm of gravity-assist trajectory and ballistic-orbit joint optimization method.Based on planar two-body gravity assist dynamics model,velocity change formulas and energy change formulas of spacecraft in central body's coordinate system after gravity assist are derived,and orbit change form is analyzed.On this basis,a judgment guideline of whether spacecraft meets next celestial body after one gravity assist is put forward.Further,based on interval analysis conception,a sequence search algorithm of non additional maneuver is proposed.By using this algorithm,all gravity assist sequences satisfying starting constraints can be found at one time,thus helping to provide selections for orbit design.After considering additional impulse maneuvers,this dissertation take Earth-Jupiter transfer as background,calculating and comparing direct transfer orbits to multi sequence gravity assist transfer orbits by using particle swarm optimization algorithm and Pork-Chop graph.Periodic window characteristics of different celestial gravity assist transfer orbits are obtained.In these results,specific points can be found that match Galileo and Juno instances.Compared with single-point result acquired by traditional orbital optimization methods,results found by method in this dissertation can provide sufficient and all-round reference for engineering mission design.Splicing calculation model considering constraints of launch azimuth and final stage parking time of launch vehicle is established.Under assumptions of non gravitational perturbation and non yawing,calculation formulas of launch azimuth and final stage parking time of launch vehicle under arbitrary deep space departure speed are derived.By using this model,variation of splicing solution range affected by departure speed,geographical location of launch site,launch azimuth and parking time are analyzed.A declination-launch energy(?-C3)map method for deep space departure matching is proposed,which can express feasible region of deep space launch capability of a given type of launch vehicle and quickly judge whether this launch vehicle is suitable for a specific deep space launch mission.After considering yawing capability of launch vehicle,this dissertation establishes a calculation model of non dimensional launch capacity under arbitrary deep space departure condition based on actual ballistic model and program.By computing departure velocity angle corresponding to transfer orbit,constraint requirements of all transfer orbits in departure window are obtained.By this means,whether an arbitrary departure condition is satisfied can be quickly computed within launch vehicle's yawing capacity,as well as its launch capacity loss.Finally,taking Earth-Vesta exploration as an example,this dissertation gives direct transfer and multiple gravity assist transfer orbit design schemes,and analyzes matching results between departure window and launch vehicle capacity,providing reference for actual mission design.