Constrained Guidance And Control Methods For Missiles With Both Aero-Fins And Lateral Jets

The dual-control technology can greatly improve the response rate of the missile control system,especially in the low dynamic-pressure environment where the control efficiency of the rudder actuator is often insufficient,and the lateral force can quickly provide the control torque,so that the missile is capable to turn rapidly.However,the jets actuators bring a variety of constraints to the control law design,including constraints on the form of the control law,constraints on the operating characteristics of the actuators,constraints on the response characteristics of the system,and constraints on the number of jets,etc.On the other hand,the missile guidance technology with the terminal constraints can improve its attack capability,for example,the constraint on the impact angle can enhance the damage effect of the warhead,and the constraint on the impact time provides the missile with the ability of salvo-attack,but for the guidance problem with both angle and time constraint,especially the three-dimensional guidance problem,there is still a lack of effective methods to solve it.In view of the current difficulties in such advanced missile guidance and control technologies,in this dissertation the research is conducted layer by layer,from local to overall,from inner-loop control to outer-loop control,and the main research contents include:First of all,taking the rotational dual-control missile as the research object,considering the factors of the missile body spin and the control characteristics of the lateral jets,a six-degree-of-freedom dynamic and kinematic model of the missile is established.On the basis of reasonable assumptions,the dual-control missile control system model is deduced and simplified as the basis for the subsequent research on missile control methods.For the jets array,aiming at improving the thrust response rate and accuracy,a rotationalsynthetic ignition logic is proposed,which lays the foundation for the implementation of the subsequent jets thrust control commands.Then,for the prescribed performance control problem of the dual-control missile,an integral sliding mode surface based on the prescribed performance is proposed,where the sliding motion is same as the referenced model with the desired performance.The missile control system uncertainty term is estimated by using the extended state observer,and a sliding mode missile control method with the uncertainty compensation is given.Based on the equivalent control concept,the adaptation law is studied,and the adaptive sliding mode control method for the dual-control missile is proposed.On the basis of the above two sliding mode control methods,considering the implementation issues,the jets command is further optimized ensuring the stability of the sliding mode surface,so that the control conservatism is reduced and the jets consumption is alleviated.In addition,for the quantized control problem of the dual-control missile,a class of second-order nonlinear systems are considered,and a novel quantized second-order sliding mode control method is given based on the referenced methods,which makes the quantized control amplitude can be reduced monotonously with the convergence of system states.It is quite simple to adjust the convergence rate of the system by adjusting quantized control parameter.Based on the concept of invariant set,the stability of the closed-loop system has been proved,where an estimation algorithm for the upper bound of the system convergence time with the worst uncertainty realization is obtained,which proves the effects of control parameter on the system convergence rate.Then the switching second-order sliding mode control law of the dual-control missile is proposed,where the jets consumption is reduced monotonously with the convergence of the system states,and the convergence rate of the missile control system can be easily adjusted.Moreover,for the guidance problem with both impact time and angle constraints,the virtual target approach is used to simplify the terminal constraints,from both the time and angle constraints to only angle constraint,and the theoretical proof is given.Based on the principle of virtual target approach,a two-stage planar guidance law is proposed.Firstly,the optimal control method is utilized to get the planar guidance law with terminal angle constraint for the virtual target in the first stage.The guidance law has an analytical form,which reduces the impact time error.Then the proportional navigation guidance law is adopted in the second stage to ensure the smoothness of the guidance law at the switching time between the two stages.Based on the analysis of the planar guidance law,some reasonable assumptions are proposed,and the three-dimensional guidance law with impact time and angle constraints is obtained,and the rationality of the assumptions is verified by the numerical simulations.Furthermore,for the multiple missiles cooperative guidance problem,based on the three-dimensional guidance method with both impact time and angle constraints of a single missile,two cooperative guidance laws,centralized and distributed,are given respectively,such that all missiles can attack the target at the same time with their own desired impact angle.Firstly,based on the leader-follower communication topology,the leader missile is set as the central node,the time-to-go of the leader missile is taken as the cooperative variable,and a centralized cooperative guidance law is proposed to make the time-to-go of the other follower missiles consistent with the leader missile.For the virtual target approach,the estimation algorithm of missile time-to-go is given.Based on the undirected connected topology,the mean of the time-to-go of each missile is taken as the cooperative variable,and the distributed cooperative guidance law is proposed to make the time-to-go of each missile converge to the mean value.Finally,based on the six-degree-of-freedom motion model of the dual-control missile,the integrated design of the missile guidance and control laws is concerned.Firstly,for the single missile guidance and control loop,the sliding mode control method based on the uncertainty compensation is selected to design its controller,and the threedimensional guidance method with both impact time and angle constraints is selected to design its guidance law.Then,for the multi-missile cooperative guidance and control loop,the centralized cooperative guidance method based on the leader-follower strategy is selected to design its cooperative guidance law.The final numerical simulation results verify the effectiveness of the missile control methods,the single missile guidance method with terminal constraints and the multi-missile cooperative guidance method proposed in this dissertation in the closed-loop cooperative guidance and control system.