Research On Optimal Midcourse And Terminal Guidance Law For Exo-atmospheric Interceptor

With the progress of space technology,the battlefield of war is no longer limited to sea,land and air,and the military activities of various countries are gradually expanding to outer space.The role of space-based military weapons is becoming more and more significant in modern wars.Therefore,the struggle for celestial power is of strategic significance to the military powers.Under this background,this thesis studies the optimization design of mid-course and terminal guidance law in space interception.Space interceptors carry limited fuel.In space interception,the flight state of the interceptor and the target aircraft is constrained by the orbit,therefore,it is necessary to design the multi-impulse mid-course guidance law optimization to reduce the fuel consumption and meet the earth radius limit when maneuve ring in a large range in space.Because of the similarity between space interception and rendezvous problem,the space interception problem is studied based on spacecraft rendezvous problem.Swarm intelligence optimization algorithm combined with prime vector theory is adopted to deal with the equality constraints by sorting time constraints.By comparing various intelligent optimization algorithms for coplanar and non-coplanar rendezvous problems,it can be concluded that differential evolution algorithm has obvious advantages in accuracy and effectiveness.When differential evolution algorithm and sorting method are used to solve a 90-deg non-coplanar rendezvous problem,it is found that it does not satisfy Lawden’s pri me vector theory.Finally,the differential evolution algorithm and the sorting method are used to solve the space intercept orbit transfer problem.The results show that the multi-impulsive orbit maneuver strategy can achieve smaller fuel consumption and satisfy the earth radius constraint condition.The combination of differential evolution algorithm and sorting method to deal with equality constraints can achieve fast and effective optimization of space multi-impulsive maneuvers.The exo-atmospheric target is flying at a high speed and the accelerati on of the interceptor is not necessarily advantageous to the target.In order to deal with the strong maneuvering of space high-speed target,a terminal guidance law with strong robustness which can work in the saturation state of guidance command for a long time is designed.A novel anti-saturation super-twisted sliding mode algorithm is proposed,which has simple structure and simple parameter setting rules,and can make full use of the limit of control input.The saturated super-twisting sliding mode guidance law is applied to the design of interceptor guidance law.The simulation verification is carried out in three space interception scenarios: when the target acceleration approaches the interceptor,the shift handover conditions at the mid-course and terminal guidance are poor,and the target acceleration exceeds the interceptor in a short time.The results show that the guidance law algorithm designed in this thesis has stronger robustness compared with the classical adaptive sliding mode guidance law.And compared with the classical high-order sliding mode guidance law,the proposed guidance law algorithm can work in saturation state for a long time.The proposed terminal guidance law has a simple structure and strong robustness,which is suitable for intercepting strongly maneuvering targets with high speed in space.In order to increase the lethality of space interceptor,a terminal guidance law considering impact angle constraint is designed.A guidance law design method for maneuvering target in two-dimensional impact angle coordinate system was proposed,and then fuel consumption was optimized by combining the opti mal sliding mode design method.Compared with the classical non-singular terminal sliding mode guidance law and the Gaussian pseudo-spectral method,it is shown that the proposed algorithm is robust and optimal when dealing with non-maneuvering or maneuvering targets.The design method proposed in this thesis is further extended to three dimensional space.The robustness and optimality of the proposed algorithm are verified in the simulation of a real missile model.The simulation results show that the three-dimensional impact angle constraint optimal sliding mode guidance law has both robustness and optimality.