Missile Guidance Laws For Intercepting High-speed Maneuvering Targets

In modern or future warfare,air threats,mainly including near-space hypersonic vehicles and maneuvering reentry ballistic missiles,have the features of high speed and maneuverability.This drives interceptor missiles to develop towards high precision,rapid maneuver,and high intercept speed.For the interception of stealth aircraft,radar dectection range of the interceptor missile shortens and so the interception of stealth aircraft imposes identical requirements on guidance law design with the interception of high-speed maneuvering target,in need of fast convergence.In above backgrounds,this thesis focuses on studying new missile guidance laws for intercepting high-speed maneuvering targets.In view of the nonlinearity of three-dimensional target-missile relative motion equation and the multiple constraints imposed on midcourse guidance phase,a three-dimensional optimal midcourse guidance law with multiple constraints,including handover range,terminal line-of-sight(LOS)angles,terminal LOS angular rates,and missile acceleration,is designed by means of Gauss pseudospectral method.In the design,the power function of time-to-go is treated as the weighting coefficient of energy cost in order to drive the missile acceleration to tend to zero at the end of midcourse guidance,contributing to a successful handover from midcourse guidance to terminal guidance.Finally,the proposed optimal midcourse guidance law is verified by simulations in the background of anti near-space hypersonic vehicles.The results show that the designated constraints can be satisfied and the convergence of the missile guidance commands to the vicinity of zero can be achieved by regulating the exponent of the energy cost weighting function.Given that first-order sliding-mode terminal guidance laws suffer from chattering and the duration of the terminal guidance process is finite,a three-dimensional smooth finite-time near-optimal integral-sliding-mode(ISM)termianl guidance law is designed by means of the combination of nonlinear non-quadratic finite-time optimal feedback control,ISM and smooth ST algorithm.Lyapunov stability analysis shows that the proposed law enables the guidance system to lie in a neighborhood of the sliding surface in the entire response process or to leave firstly and then return to the neighborhood in finite time,and ultimately enables the target-missile tangential relative velocities to converge to zero or its neighborhood in finite time.Smooth ST algorithm guarantees that the proposed law is not only continuous but also smooth and so chattering-free,apt to be applied in practice.Compared with the conventional optimal guidance laws,it does not require time-to-go estimation,apt to be implemented.Finally,the performances of finite-time fast convergence,strong robustness,near optimality,and chattering rejection of the proposed near-optimal ISM terminal guidance law are confirmed via simulations in the background of anti near-space hypersonic vehicles.Also,high guidance precision is yielded.In consideration of the adverse effect of missile autopilot dynamics on the gudiance performance,the research on finite-time guidance laws should be extended to consider autopilot dynamics.A finite-time disturbance observer to estimate target maneuver information is designed and the estimates of the matched and mismatched disturbances resulting from target maneuver are given.In order to reject the mismatched disturbances,the modified states are defined based on the estimates of the mismatched disturbances.The modified state-related nonlinear integral sliding surface is defined and a finite-time ISM terminal guidance law with second-order autopilot dynamics is designed by means of smooth ST algorithm and the estimate of the matched disturbance.It is proved that the proposed law can not only force the sliding mode to occur in finite time but also drive the target-missile tangential relative velocities to converge to zero in finite time using finite-time bounded function and Lyapunov approach.Moreover,the proposed law requires neither the derivatives of the LOS angular rate nor the target acceleration,apt to be applied in reality.Simulation results in the background of anti stealth aircraft and anti reentry ballistic missile exhibit that the proposed law is robust to target maneuvers and variations of initial engagement conditions.It can achieve finite-time fast convergence,effectually make up for the impact of autopilot dynamics,and reject the mismatched disturbances.It yields high guidance precision even in the presence of seeker measurement noise.Considering that the available acceleration of a missile must be finite in practice,it is essential to design gudiance laws with acceleration saturation constraint.A nonlinear auxiliary design system ensuring finite-time convergence is designed to analyze the effect of acceleration saturation constraint.Compensated state is defined in virtue of auxiliary design system state.On the basis of compensated state dynamics,a smooth finite-time near-optimal ISM terminal guidance law with acceleration saturation constraint is synthesized by combining finite-time optimal feedback control,ISM,and smooth ST algorithm.Its properties of finite-time convergence and robustness are demonstrated by means of Lyapunov stability method.Simulation results in the background of anti stealth aircraft and anti reentry ballistic missile show that the proposed law can handle saturation constraint effectively and possesses high guidance precision.Moreover,a three-dimensional nonlinear command filtered backstepping terminal guidance law is proposed in order to handle acceleration saturation constraint while making up for autopilot dynamics.Command filters avoid the phenomenon of“explosion of terms”and the introduction of immeasurable feedback quantities caused by computing analytic derivatives of virtual controls.In the design,several auxiliary design systems are introduced to analyze the effect of command filtered errors and accelertion saturation constraint in order to compensate for them.It is proved that the closed-loop guidance system under the command filtered backstepping guidance law is semi-globally uniformly ultimately bounded.Simulation results confirm the effectiveness of the porposed law in handling acceleration saturation constraint and compensating for the effect of autopilot dynamics.Its fine robustness against maneuvering targets and high guidance precision are also demonstrated.For intercepting unpredictedly maneuvering targets,the missile guidance problem with second-order autopilot dynamics is formulated into disturbance attenuation linear output feedback H_?control problem by means of the idea of feedback linearization.This control problem is ultimately transformed into two LMI constraints.The proposed law is only in need of target-missile relative range,target-missile closing velocity,and line-of-sight angular rate and it does not require target acceleration and its bound,apt to be implemented.Finally,the proposed law is verified in terms of anti unmanned aerial vehicle(UAV)and anti reentry ballistic missile.The results exhibit that it is robust to target maneuvers and variations of initial engagement conditions,it can make up for the impact of autopilot dynamics effectively,and it yields high guidance precision.