Research On The Guidance Law Design For Defensing The Hypersonic Target In Cruise Phase

Near space hypersonic vehicle has the characteristics of high speed and strong maneuverability,which makes it an offensive weapon with strong development potential.Traditional air defense anti-missile systems are difficult to effectively defend against near space hypersonic targets.In view of the characteristics that this kind of target is easily detected and tracked when it is in the cruise section,the paper studies the defense guidance technology for this hypersonic cruise target.Taking the air base interceptor as the research object,aiming at the operation mode and the whole course guidance process of the medium and long range missile system,the guidance law of the single interceptor unpowered flight section under the condition of middle and terminal guidance under the condition of small speed ratio,the multi interception missile in the non dynamic flight section under the condition of small speed ratio,and the coordination guidance law of multiple interceptor projectiles and a variety of complexity are studied.From the two aspects of seeker and trajectory hand-over,the mid and terminal guidance cross over problem of high-speed target is analyzed and studied.For the hand-over problem of the seeker,the range of the field of view of the capture area and the starting time of the seeker at the time of the seeker is combined,and the effective hand-over area of the interceptor at the time of the hand-over is given.On the other hand,in order to ensure the smooth transition of the two guidance segments and avoid the jump of the two segment trajectory near the intersection point,a medium and terminal guidance method based on Hermite interpolation is proposed by the maximum allowment course error.Taking the mid-course and terminal guidance transfer conditions as the terminal constraint conditions of the mid-course guidance,the unpowered guidance flight law of the engine after exhaustion of fuel is studied.Combined with the model predictive static programming theory,the analytic algorithm of the optimal control energy consumption problem is derived and used as an instruction revision guidance law for interceptors.The guidance law based on trajectory optimization is very sensitive to initial guess,and another kind of midcourse guidance law based on inscribed angle is designed.By using the method of non singular terminal sliding mode control and extended state observer,the disturbance control amount is designed to compensate the influence of the disturbance part in the dynamic model of the inscribed angle.A cooperative problem in mid-course guidance phase is addressed in this paper.For providing suitable initial conditions of successful terminal salvo attack,a novel finite-time cooperative mid-course guidance law with terminal handover constraints is proposed.Firstly,a three-dimensional guidance model of mid-course is decoupled in a planar line-of-sight frame as a two-point boundary value problem.The terminal handover constraints which can guarantee an ideal zero effort terminal engagement are proposed and analyzed.Secondly,the design of the cooperative mid-course guidance law is separated into two stages.The acceleration commands along the line-of-sight direction are developed based on the finite-time average-consensus protocol and super-twisting algorithm in the first stage.In the second stage,the model predictive static programming method is adopted to solve the two-point boundary value problem in LOS frame with a known final time from first stage.Furthermore,sliding mode control theory is used in combination with model predictive static programming method to satisfy terminal handover constraints with bounded perturbation.A two-phased guidance problem with terminal impact angle constraints and seeker's field-of-view limit is addressed in this paper for a missile against a non-maneuvering incoming target.To avoid the look angle exceeding the seeker's physical limit,a two-phased look angle control guidance scheme with the terminal constraint is introduced.The PN-typed guidance law is designed for each guidance phase with a specific switching condition of line-of-sight.The proposed guidance law is comprised of two types of acceleration commands: the one in the initial phase which aims at controlling the missile's look angle to reach the limit,and the other for final phase which is produced by switching the navigation gain.The monotonicity of the line-of-sight angle and look angle is analyzed and proved to support the proposed method.In addition to the conventional rudder deflection control instruction limit,the dynamic characteristics of the rudder deflection angle will also be limited by physical factors.The rotation angle rate of rudder deflection corresponds to the change rate of command acceleration in guidance system.Aiming at this problem,the weighted optimal guidance law with attack angle constraint and instruction acceleration variation rate constraint is studied in this paper.An optimal guidance law is designed by using the first derivative of the missile acceleration as the control variable instead of the traditional acceleration.According to the weighted optimal solution of the linear system,the guidance law is designed to achieve the constraint of acceleration change rate by weighting function.