Research On High-precision And High-efficiency Terminal Guidance And Control Key Technologies For Missile

In this dissertation, the rearch background is the advanced control and precise guidance technique of missile and this work is focused on design and analysis of terminal guidance and control system of modern homing missile. In close connection with research status and development trend both at home and abroad about Multi-Input and Multi-Output stability margin analysis, missile guidance laws with terminal constraints as well as integrated guidance and control, some key technologies for missiles, such as control design and singular perturbation margin analysis, optimal guidance laws for attacking maneuvering targets, nonlinear guidance schemes with terminal constraints, integrated guidance and control and so on, are systematicly studied and thoroughly exploited. Thereafter, several main research subjects are involved:Firstly, research background, significance, and current difficulties are discussed and, application status and development trend of domestic and oversea typical guided weapons are introduced. Some difficulties of key enginnering technique including stability analysis of control system, guidance laws with some terminal constraints, integrated guidance and control design and so on, are specifically summarized. At last, research scheme, content structure, and innovation points are given for this work.Secondly, definitions and tranformations of some common coordinate systems are given and the complete differential equations of missile motion are built for endoatmosphere missiles. Mathematical formulas solving for some important angles are derived and expressions of aerodynamic forces and monments as well as how to use datum from wind tunnel test are also stated. The configuration of missile guidance and control system is addressed briefly.Next, the concept of singular perturbation margin for Multi-Input and Multi-Output systems is proposed to measure the robustness of missile autopilot and then the method of calculating this stability margin is given to provide a specific metric. Pole assignment design method for three-loop autopilot and LQG/LTR design technique for integrator lateral autopilot(that is, turn coordinate control system) are addressed. The proposed metric has been successfully applied to estimate the robustness of previous two autopilots.Under consideration of the lag property of autopilot, optimal control theory is used to derive guidance laws with compensation of complicated dynamics of autopilot for attacking constant and weaving maneuver targets respectively. This approach consisits of the numerical integration of the nonlinear Ricatti differential equation to solve for the guidance law control gains. Kalman filter and Extended Kalman filter for two types of target maneaver are respectively proposed to eastimate target properties online. Simulation results have demonstrated that the combined guidance scheme of the proposed optimal guidance law and corresponding Kalman filter effectively compensates the miss distance resulting from target maneuver, autopilot lag property, and ‘wrong-direction’ effect.In order to improve lethality, the guidance problem with impact angle constrainted flight trajectories is dicussed using finite-time control theory. First, double-integrator finite-time stablization control law is generalized. The continuous finite-time stabilization guidance law is proposed in conjuction with feedback linearization technique. This guidance law is analyzed by the finite-time stabilization homogeneous theory, thus confirming the globally finite-time stable property. Based on the finite-time Lyapunov stability theory, the expression of convergence time is yielded. The proposed guidance law ensures to shape the missile trajectory near impact in advance, that is, to make the line-of-sight rate as well as line-of-sight angle error approach zero as soon as possible.Finally, the terminal guidance problem with attack angle constraint that accounts for the lag of autopilot is discussed in order to improve lethality further. The generalized integral back-stepping theory is proposed and proved, and then back-stepping guidance law without consideration of the lag of autopilot is designed. Furthermore, the guidance law with impact angle constraint for the case of the lag of autopilot is derived utilizing the back-stepping method again, and the exponentially convergent characteristics of the line-of-sight angle and angular rate for the proposed guidance law is detailedly proved using Lyapunov stability theory. The proposed guidance law is successfully applied to six-degrees-of-freedom mathematical simulation for some guided weapon, achieving smaller miss distance and saving a large part of energy.The work of this dissertation provides some resolutions for the design and analysis of missile control system and terminal guidance law. Those schemes are demonstrated to be effective and feasible through theory analysis and mathematical simulation.