Time-delay Problem Study And Fault-tolerant Control Design Of Aircraft Wing Flutter

In order to enhance the flight performance,modern aircraft designers pay more attention to the improvement of aerodynamic efficiency and the decrease of structure weight,the structure of aircraft is designed to be more flexible,which leads to the aeroelasticity problem.Since aeroelasticity is a multidisciplinary problem with strong nonlinearity and uncertainty,it has always been the emphasis and difficulty in aircraft design.The study of aircraft aeroelasticity problem is of great scientific significance and engineering value.Missile is a member of the aircraft,and its slenderness ratio has increased.Large flexible deformation will couple with aerodynamic force and control force,which will have an effect on the stability and accuracy of control of missile,as well as fight performance.How to design the guidance and control system for missile to ensure the precise hit is worth studying.The flutter phenomenon is one of the most important problems in aeroelasticity.When wind speed is larger than the critical flutter speed,flexible force of the structure,inertial force and aerodynamic force will couple under unsteady aerodynamic pressure,creating wing flutter.If the flutter is not suppressed in time,the vibration will continue to grow,ultimately leading to aeroelastic instability or catastrophic structural failure.This research was funded by The National Natural Science Foundation of China(Grant Nos.11132001,11272202),the Key Scientific Project of Shanghai Municipal Education Commission(Grant No.14ZZ021)and the Natural Science Foundation of Shanghai(Grant No.14ZR1421000).Guidance and control system of missile,wing flutter model of aircraft,as well as the time-delay control and fault-tolerant control of wing flutter are studied extensively in this dissertation.The main research and achievements are as follows:(1)The guidance and control system of missile is designed.Firstly,the mathematical model of the missile is established based on the dynamic principles.Then,based on the established model,the guidance law is studied using the generalized proportional navigation law.Thereafter,the missile autopilot is designed using the traditional three channel design method.Meanwhile,considering the coupling of missile's flexible vibration and aerodynamics,the vibration equations with the coupling terms are also established.The reserch results show that:(i)considering axial speed,the damping in dynamic equation of flexible missile is generated,and the stiffness in the equation is changed.The system tends to be more unstable as the slender ratio increases,and the responses become larger.If the axial speed increases,the stiffness will reduce and the system will become more unstable with larger responses.(ii)guidance law designed by proportional navigation and the three-channel independent control system join with coordinated circuit is effective.Every movement parameters well follow the control instruction and the missile meets the requirements of hitting maneuvering targets.(2)The flutter model of two-dimensional wing is established.On the premise of subsonic,the Theodorsen unsteady aerodynamic model is firstly introduced.Using the Lagrange method,the dynamic equation of airfoil flutter is established,in which the cubic hardening spring nonlinearity of pitch stiffness is considered.Finally,the critical flutter speed of the airfoil is acquired using the V-g method.The simulation results show that:(i)without considering nonlinear stiffness,the system is stable when the wind speed is less than the critical flutter speed;the responses of the system generate harmonic motion when the wind speed is equal to the critical flutter speed;the system is unstable when the wind speed is larger than the critical flutter speed.(ii)if the nonlinear stiffness is taken into account,the plunge displacement and pitch angle will appear stable periodic solution near the origin,and generate limited circle oscillation.(3)The time-delayed feedback control of wing flutter is studied.The state equation with time delay is transformed into a standard state equation with implicit time delay by a special integral transformation.Two nonlinear time-delay controllers are designed using the instantaneous optimal control method and the sliding mode control(SMC)method respectively.The simulation results indicate that:(i)time delay in the control system has significant influence on the control performance.Control failure may happen if time delay is not considered in control design.(ii)The time-delay controller proposed are effective in suppressing the wing flutter and the SMC one can suppress the flutter with either small or large control time delay.(4)The fault-tolerant control is studied.Under the circumstance of partial loss of actuator effectiveness performance,two situations are discussed.Firstly,using on-line updating law to estimate the bound of actuator fault,an adaptive sliding mode fault tolerant controller is derived without considering input saturation.Next,considering input saturation,an input constraints error dynamic enlargement method is adopted in order to settle the problem of actuator saturation;combined with the fault-tolerant controller proposed before,the final control law is acquired.The simulation results indicate that the adaptive sliding mode fault tolerant controller is effective in suppressing the airfoil flutter under circumstance of partial loss of actuator effectiveness performance and input saturation.