Research On Guidance And Control Of High Speed Aircraft Based On Sliding Mode Method

With the development of aircraft technology and the improvement of national defense demand,high-speed aircraft,as a new type of weapon,has received close attention from various countries.Guidance and attitude control system is two important parts of high-speed aircraft control system.The flight environment of high-speed aircraft is complex,in order to ensure the accurate and reliable completion of the flight mission,it is necessary to carry out the research on the guidance law and attitude control law of the aircraft.Firstly,the common coordinate system used to describe the motion of the aircraft and the conversion methods between different coordinate systems are introduced.Under certain assumptions,a nonlinear model is established for the attitude motion of the high-speed aircraft based on the dynamic principle and the kinematics principle,which lays a foundation for the subsequent chapters..Secondly,the guidance law based on sliding mode method is studied.Firstly,in order to reduce the chattering of the traditional sliding mode controller,an adaptive sliding mode method based on particle swarm optimization is proposed.In order to complete the target interception task,the guidance law with attack angle constraint is designed according to the sliding mode method,and the capture ability of the designed guidance law is analyzed.The off-target condition is analyzed and the original guidance law is improved according to the off-target condition.Numerical experiments show that the improved guidance law can meet the attack angle constraint and avoid off-target situations.In order to meet the attack time constraint,a time-constrained sliding mode guidance law is designed on the premise that the flight time estimation is accurate.The simulation proves that the proposed guidance law can accomplish the task of many-to-one coordinated attack.Then,in order to meet the needs of guidance system,an attitude control law of the aircraft based on adaptive terminal sliding mode is proposed.According to the multi-time scale principle,the attitude control system of the aircraft is divided into two loops,the outer loop is the attitude angle control loop,and the inner loop is the attitude angular speed control loop.The outer loop generates the virtual angular speed command as the input command of the inner loop.The terminal sliding mode control laws are designed for two loops.It is proved that the design control law is non-singular.In order to improve the robustness of the system and weaken the chattering of the sliding mode controller,a nonlinear disturbance observer with finite time convergence is designed to observe the disturbance in the system,and the estimated value of the disturbance is used in the control law.The effectiveness of the disturbance observer and controller is verified by numerical simulation.Then,in order to ensure the robustness of the system and weaken the chattering of the sliding mode controller,a quasi-continuous sliding mode attitude control law which increases the relative order of the system is proposed.The attitude control system of the aircraft is divided into two loops.The full-order sliding mode method is adopt in the outer loop to design the sliding mode control law,and first order inertial link is used to obtain smooth attitude angular speed commands.A quasi-continuous high-order sliding mode control law is designed for inner loop by introducing virtual control quantities and increasing the relative order of the system.In the case of existing parameter uncertainty and external disturbance,the proposed method is used to carry out numerical experiments,which proves that the algorithm has strong robustness.Finally,the paper summarizes the work of this paper and analyzes the problems still existing in the aircraft guidance law and attitude control law,which provides direction for future research.