Study Of Deorbit And Reentry Trajectory Design And Guidance Methods For Lifting Body Spacecraft

Lifting body spacecraft is a type of hypersonic glide-reentry vehicle, which can deorbit from near Earth orbit in a very fast way, and thus will be an important part of the future Earth orbital transportation system. Focusing on lifting body spacecraft’s return process from various Earth orbits, the associated trajectory design and guidance problems are studied. The main work and achievements are as follows:The deorbit window design for lifting body spacecraft is studied. Firstly, the effects of the maneuverable downrange and the cross range on the reentry window are analyzed; then, an orbit phasing approach is adopted to speed up the spacecraft’s falling into the reentry window; lastly, based on the above research, assuming the initial orbit is a circular orbit of 400 km altitude and characteristic velocity is less than 150m/s, an orbit maneuver strategy is designed. Based on large sample of initial position, simulation results indicate that 77.06% of the sample’s waiting time are less than 24 h, and 22.94% of the sample’s waiting time are between 24 h and 43.27 h.A single “burn-coast” deorbit trajectory optimization method and a “multi-burn-coast” deorbit trajectory planning method are presented. The idea of the former method is a combination of direct method and indirect method, which uses a combination of ES algorithm and SQP method to identify the value of the costate variables, and to solve the two point boundary value problem. By the idea that the energy and angular momentum should decrease synchronously at approximately the same relative rate, a combinational deorbit trajectory planning approach is therefore proposed, which consists of tangential thrust firing during braking phase and a characteristic-function-based criterion for the engine cut-off logic. Simulations of different altitude orbits proved the effectiveness of the proposed two methods.The existence of a critical altitude in the single pulse deorbit problem is proved, and the conclusion is extended to the finite thrust deorbit problems. By the theory of solution existence, the proposition that the critical altitude value depends only on the entry interface conditions is proved. When the initial orbit is lower than the critical altitude value, the velocity-gain guidance works well. To address the deorbit problem of solid rocket-powered spacecraft, a depleted shutdown hybrid guidance method including energy dissipation and closed-loop guidance is presented. According to entry interface conditions,the models of surplus apparent velocity module prediction and energy management are given based on standard thrust assumption. Given continuous thrust, the features of spacecraft energy and angular momentum are analyzed, and the concept of energy window is presented, following which, a random surplus fuel dissipation approach is proposed via switching the direction of the angle of attack to achieve energy control.In order to solve the problem in which the initial orbit is higher than critical altitude value, a "burn-coast-burn-coast" deorbit guidance approach, based on energy and angular momentum indices, is presented. The trajectory is supposed to a "burn-coast-burn-coast" style, with the thrust direction of the first burn opposite to the velocity direction, and the second burn using the guidance law. Deorbit simulations of different initial orbit altitudes and interface conditions show that the method can effectively solve the deorbit guidance problem which the single "burn-coast" method fails to do, with less computational efforts.The optimization methods for the initial entry phase and equilibrium glide of the spacecraft are studied. Due to the requirements of the engineering, the entry trajectory is divided into two phases; an initial entry one and an equilibrium glide one. Different direct optimization methods are applied to these two phases. The attitude maneuver ability of the spacecraft is poor during the first phase, so the associated attack angle and bank angle are supposed to be constant during each discrete time interval. By optimizing these unknown variables, fast planning of the trajectory can be realized. For the second phase, the hp-adaptive pseudospectral method is adopted. The method can adjust the number of segments and the degrees of polynomials in each segment in an adaptive way.The guidance approaches for the initial entry phase and the equilibrium glide phase are studied. Due to the poor maneuver ability of the spacecraft during the first phase, traditional D-V profile following method can well satisfy the constraints on the terminal velocity, altitude and flight path angle. The maneuver ability of lifting body spacecraft during the second phase is much enhanced. The proposed guidance approach based on LQR theory can be well suited to the complex entry environment. Simulations show that the LQR method has good performance in range following, velocity following and flight path angle following.This dissertation expands the research domain of the trajectory planning and guidance for lifting body spacecraft. Owing to the idea of combining theory with engineering practice, some new trajectory design and guidance approaches during deorbit phase and entry phase are developed. The results may be a good reference to the overall design of lifting body spacecraft.