Mars Aerocapture Trajectory Design And Guidance

A spacecraft can enter an orbit around the targeted celestial body fast,precisely and fuel-saving with aerocapture.It effectively resolves the contradiction between fuel consumption and time cost in the design of capture orbit,thus it has a good prospect of application.This work was supported by the National Basic Research Program of China:Research on navigation,guidance,and control problem for planetary precision landing(No.2012CB720000).In the background of Mars exploration mission,this dissertation concentrates on methods for aerocapture trajectory design and guidance.The main contributions of the research are summarized as follows:For Mars aerocapture dynamic modeling problem,the Mars environment model is first provided systematically.The common simplified models of Mars atmospheric density are introduced,classified,compared and analysed.The aerocapture dynamics is introduced,including orbital mechanics and flight mechanics.For aerocapture trajectory optimisation and guidance design,models in different guidance methods are introduced in detail.This part of research is the foundation for subsequent parameter determination of target orbit,the identification of entry and exit conditions,the guidance design,simulations,and evaluations in this dissertation.The analysis and determination methods of aerocapture target orbit parameters with near minimum fuel consumption are presented.For the difficult problem of deceleration in direct entry,descend and landing(EDL)resulting from high entry velocity,large aerodynamic load,large heat flux and so on,the scheme is proposed that a spacecraft is captured into a target parking orbit by aerocapture and then enter the atmosphere.To reveal the relationship between fuel consumption and parking orbit parameters in the scheme,the determination method of aerocapture target orbit parameters with minimum fuel consumption for EDL is developed.The analytic expression and its partial derivative of velocity increment are derived,dealing with the manoeuvre that transfers the post-aerocapture orbit into the target orbit.The condition for minimum fuel consumption is determined by judging monotonicity of the partial derivative.The expression of velocity increment of the manoeuvre that transfers the target orbit into the deorbit are derived.For the complexity of the expression,a method based on number-shape combination is presented to analyse the condition for minimum fuel consumption by contour maps.Combining the above two conditions,the condition for minimization of total velocity increment is provided by Gaussian pseudo-spectral method(GPM)that obtains the relation between target orbit parameters and fuel consumption.Numerical computation results show that the apse line and apoapsis of the target orbit are the same with those of the deorbit,and the periapsis of the target orbit should be as low as possible to reduce the fuel consumption.The exit conditions for aerocapture with minimum fuel consumption are given,and the method for searching entry controllable set corresponding to the exit conditions are presented.For the fuel consumption for entering target orbit by aerocapture,the condition for minimum fuel consumption when transferring the post-aerocapture orbit into the target orbit is analysed.The relation between orbital elements of post-aerocapture orbit and that of target orbit and exit flight path angle are derived.The domain of entry velocity corresponding to exit flight path angle is analysed by aerocapture dynamics with the consideration of path constraints.The exit flight path angle satisfying both the requirement of the domain of entry velocity and fuel-saving is obtained,and then the nominal exit condition is provided.Based on the theory of controllable set,an approach is presented to compute the entry condition for maximum entry zone according to the nominal exit condition.The trajectory satisfying the entry and exit conditons for aerocapture is obtained by GPM,which demonstrates the correctness of the entry and exit conditions.With the above trajectory as nominal trajectory,the exit states and post-aerocapture orbit under non-perturbation conditon and Monte Carlo conditions are obtained by open control with bank angle profile of nominal trajectory.Numerical computation results show that the smaller the exit flight path angle is,the less the fuel consumes,but the selection of it should consider the actual domain of entry velocity,and the higher the entry velocity is,the larger the domain of entry flight path angle and entry zone are.The results with open control under non-perturbation conditon have high precision,while have low precision under perturbation conditon.The aerocapture numerical predictor-corrector guidance with exit angular momentum as performance index is proposed.Since the frequently used methods in literature that the apoapsis altitude as the target parameter for guidance cannot ensure the minimization of fuel consumption for transferring the post-aerocapture orbit into the target orbit,exit angular momentum for guidance are presented.The heading angle are presented as the bounds of bank reversal logics for lateral guidance design.The scheme design of aerocapture guidance and numerical predictor-corrector guidance are elaborated respectively.The model for designing aerocapture numerical predictor-corrector guidance is established.The guidance results of exit angular momentum and apoapsis altitude are analysed under non-perturbation condition and conditions with entry deviation,spacecraft parameter deviation and environment perturbation.The target parameter for guidance with fewer fuel consumption for orbital transfer from the post-aerocapture orbit into the target orbit is obtained.Numerical computation results show that the exit angular momentum as the target parameter for guidance can save more fuel for orbital transfer from the post-aerocapture orbit into the target orbit compared with apoapsis altitude as the target parameter.The lateral parameters of results with different target parameters for guidance are relatively close,and their deviations are relatively small,and the velocity increment for orbital plane angle correction are very close.The results show that the heading angle as the bounds of bank reversal logic for lateral guidance is applicable to different target parameters with high precision.