Research On Flight Attitude Control Method Of Cruise Missile

Cruise missile,an important strategic weapon in the modern battlefield,has the characteristics of high-precision strike,high reliability and good cost performance.Therefore,many countries have laid more and more emphasis on their research of cruise missiles.Both excellent guidance system and attitude control system leads to the high-precision strike of cruise missiles.There has been plenty of research on guidance system and attitude control.The main research purpose of this thesis is attitude control.This thesis designs the control system for the cruise flight section and terminal maneuver section of the cruise missile.Its advantages and disadvantages are verified through MATLAB simulation.Firstly,the mathematical model of cruise missile is established.The actuating unit is aerodynamic rudder.Both the kinematics and dynamics model of the cruise missile is established by the force analysis and coordinate transformation principle,then the attitude control model of the missile is derived.After simplifying the above model with three-channel decoupling,the attitude model of the pitch direction is obtained,which lays the foundation for the subsequent controller design.Secondly,the attitude model is further simplified by coefficient freezing and small perturbation linearization.The linearized attitude model is used as the controlled plant of the cruise flight segment of the missile.The flight parameters of the controlled plant in the cruise flight segment will evolve over time.Here five characteristic operating points are selected.The designed control system that should have strong robustness and anti-interference ability,requires the missile to fly stably under time-varying parameters.Two control methods mainly used in this thesis will be introduced as follow:The first method is PID control.First,the stability of the projectile is improved by proportional feedback and series lead correction,and then the PID controller of the inner and outer loops is designed to achieve good tracking performance and anti-interference performance.The same design method is adopted for the five characteristic points,and the continuous controller of the whole process is obtained by adopting the parameter linear time-varying method between the characteristic points.The second method is the Quantitative Feedback Theory（QFT）.For the defects in the PID control system,such as time-varying parameters,double closed-loop structure,and difficult to strictly meet the performance,In order to solve the problem that traditional QFT controls the object for a long time to suppress interference,a QFT based on feedback correction is proposed.By designing feedback correction links,the boundary of anti-interference performance is broadened and the accessibility of controller design indicators is improved.In the design of the QFT controller,the variation range of the parameters of the controlled object is first obtained according to the parameters of the five characteristic points,and three performance indicators of robustness,anti-interference and tracking performance are established,and the corresponding parameters are obtained by the method of quadratic inequality.The simulation results show that the control system can strictly meet the performance indicators in the entire flight segment,and it is still robust to large parameter mismatches.Then,the direct side-jet force control model is introduced into the missile attitude control model,and the compound force missile attitude control model is obtained.As the controlled object of the terminal maneuver flight segment,the maneuver flight segment has the characteristics of large parameter changes and unknown interference.The design requires that the missile can complete a large attitude adjustment in a short time,with rapidity and stability.Its control method and control amount distribution method are mainly researched:Sliding mode variable structure control is adopted in the control method.In order to suppress the chattering phenomenon of control variables in traditional sliding mode control,the extended state observer（ESO）in active disturbance rejection technology is introduced to design the disturbance observation feedback link.According to the Lyapunov stability criterion,the convergence of the control law is proved.Finally,the tracking performance of the control system,the reachability of the control quantity and the robustness of the ESO are verified by simulation.In terms of the distribution method of the control amount,the proportional distribution method and the chain distribution method are the main methods used.In the proportional distribution method,in order to solve the problem of large deviation of the control quantity in the traditional proportional distribution method,a proportional distribution method based on the tracking deviation and the control quantity distribution deviation is proposed;in the chain distribution method,considering the angle saturation and The angular velocity is saturated,a chain allocation method based on rudder deflection constraint is proposed,and the effectiveness of each is verified by simulation.Finally,the aerodynamic efficiency of several allocation methods is compared in terms of adjustment time,overshoot,and fuel consumption.The performance of large and small instructions is tracked in the environment,and the applicable conditions of several allocation methods are summarized.