The Variable Structure Control Of Cruise Missile

The attitude control of Cruise Missile System is a highly nonlinear time-varying system, the main difficulty of attitude control derived from the uncertainty of dynamic parameters. The traditional Missile Control System largely utilized the combination technology of fixed gain and feedback gain, namely PID controller. It is widely used at one time that PID controller is applied successfully in aviation and spaceflight, but the effect of PID is not always satisfied along with the increasing demand of precision. The experts and engineers in missile control try their best to apply advanced control theory to the design of Missile Control System with the development of control theory, which could reach the requirement of modern war. This paper did some research on the attitude control scheme of Cruise Missile System based on a Cruise Missile.First of all, it is briefly introduced that the signification and present status of attitude control of missile. Then the mathematical model of missile, which is the base of analysis, design and simulation, is established. The dynamics of missile is simply considered as a rigid body dynamics in space, which is the combination of translation and rotation. The translation is researched by Newton's Law, and the rotation is researched using Moment of Momentum Theorem.In this paper, the missile attitude control is studied; especially the movement around the center of mass is analyzed. The missile dynamic equations of motion around the center of mass rotation and missile dynamic equations are given through the force analysis and dynamic center of mass theorem. And then combined with the conversion of space coordinates; we get six DOF missile dynamic equations and kinematics equations. The equations are obtained by multiple nonlinear equations, to facilitate the research; these series of nonlinear equations are transformed into linear equations. Based on income in these equations are simplified to obtain the small disturbance equation in vertical channel group, which is considered as the mathematical model for this control method.The general characteristics of the missile attitude control are: time-varying motion parameters, the serious non-linear control object, random variability of parameters and disturbances. The time variation of Motion parameters is an important feature of the missile attitude motion, which is the fuel consumption, thrust change, trajectory change factors. The variability of parameter brings two problems for the controller design of the attitude system that are how to deal with the variable coefficient and how to adapt the variable parameter. The Frozen Coefficient Method is the common method in the missile attitude control system design. The basic idea of the Frozen Coefficient Method is that it needs pre-select some moment to "freeze" the system parameters and in these moment the system can be treated as a linear system for analysis and synthesis.For the serious non-linear control object problem, the actual line of assumptions are often obtained under the design of attitude control system, then the nonlinear system for a large number of digital simulation to test attitude control system can be appropriate for some of the controller optimal adjustment of parameters.As for the random variability of parameters and disturbances. Usually deal with these random disturbances into the deterministic interference. Specific approach is to take the maximum value of random disturbances, according to the most unfavorable combination of interference simulation, ensure that the system has sufficient anti-jamming capability.For missile attitude control characteristics of the above, in this paper, we design a variable structure control approach. Use the variable structure control deal with system with uncertainty is a very effective way. the variable structure control approach has many advantages, for example parameter perturbation and interference invariance, break down and reduce the order of the system, algorithm is simple and easy to implement, etc. For missile, as the flying altitude, speed and weather conditions change will lead to diversification of the dynamic characteristics, this creates a projectile parameters are time-varying. Since a notable characteristic of the variable structure control is to parameter variation and external disturbances have strong robustness in the sliding surface, therefore we can effectively solve the missile's attitude control problem. But variable structure control also have serious flaws. Actual system because the existence of the inevitable inertia switching device, variable Structure Control in different control logic switch back and forth, Cause actual sliding movement is not accurate in switching surface, can trigger severe system jitter, making its application in the real system a major obstacle.According to the missile flight in a large airspace, the system parameters have a range of changes around the nominal value bring uncertainty problems to the system. This article points on this issue to design a variable structure controller. However, the actual missile flight, the uncertainty is difficult to determine its upper bound. The process of the missile in flight subject to the thrust line deflection and as the center of mass traversing to bring structural interference, wind interference random noise.For simplicity, the common practice is to change the random distortion and parameters uncertainty into certain distortion. In detail, it is done by taking the maximum of the random distortion and uncertainty, and then proceeds into system simulation using the most disadvantageous combination. However, due to the controller, the higher the f, the more vibration the system is. And over-vibration causes problem to the stability of the system, even causing the system to diffuse. In order to solve this problem, this report estimates the f using fuzzy logic method and adaptive method respectively, thus designing two variable structure controller.During experiments, the two controllers both well solve the problem of parameter uncertainty and random distortion.The focus of this report is to analyze the missile during its flight, kinematics model. Due to the fact that the missile control system is complicated as 6 degrees of freedom, this report will only discuss how the missile behaves under small distortion. According to the characteristics of the missile control, the kinematics model should first be linearised, then simplified into final result. Based on the longitudinal channel profile control method, the design has been proposed for the two controls on variable structure control method. As the characteristics of the missile and complexity of the attitude control system, the method in this report is only a preliminary study of rigid body movement of missiles designed longitudinal channels. However for the missile attitude control of flexible movement of the impact of multi-channel and 6 degrees of freedom attitude control problem, research and study will be further carried out in the future.