Research On Analysis Of Propagation Mechanism And Compensation Method For Positioning Orientation Error Of Ballistic Missile

Range ballistic missile is the main force of China’s strategic deterrence,therefore, to improve the hit accuracy of mobile launch missile is very urgent. Aswe all known, the Earth is an extremely complex irregular sphere, which causedthe initial positioning and orientation error is the inherent systematic bias.Thence, this paper aims to study propagation mechanism of the positioning andorientation errors, and its impact on the accuracy of iterative guidance, and itscompensation method adopted by staller-Inertia composite guidance. The mainwork of the thesis includes:Firstly, in order to quickly calculate the placement information of the missile,differential equations of motion are established in a non-orthogonal absolutecoordinate system when begins free flight, then the analytical solution is solvedthrough the transform of the relevant variables, so the orbital parameters at anypoint can be calculated by the given angle of range. Finally, the effectiveness ofanalytical solution of free flight trajectory is validated by simulation, whichmade the calculation time is greatly reduced.Secondly, supposed the positioning and orientation errors are smallperturbations, the navigation perturbed equation based on state spaceperturbation method is established, then the error model of apparent accelerationand initial velocity and analytical expressions of gravitational partial derivativematrix are deduced, and the analytical solutions of navigation perturbed equationand the spreading weight matrices are derived. Finally, the state deviations atshutdown point and the placement deviations can be simulated by analyticalsolutions, from the simulation results, the analytical solutions are consistent withactual deviations, which verifies the validity of the method.Thirdly, the optimal guidance model outside the atmosphere is established bymaximum principle under the influence of the positioning and orientation errors,then converts the terminal constraints into a group of easily guaranteed constraintvectors, the optimal thrust direction in each guidance period is derived by theemployment of Newton iteration with the help of the performance indicators forthe shortest flight time and the iteration variables for conjugate vectors and the third-stage flight time. The simulation shows that the iterative guidance used inthis paper can better suppress the influence of the initial positioning andorientation errors.Finally, after the end of the boost stage, the misalignment angle of the platformcoordinate system relative to the emission inertial coordinate system can beestimated by the adoption of dual vector attitude with the help of star sensormeasurement information, which can derive the value of the initial positioningand orientation errors. The estimation of state deviations at the shutdown pointand the placement deviations are obtained by different combinations ofpositioning and orientation errors. The simulation indicates that the estimationaccuracy accords with actual deviations, and the amended CEP reduces fromnearly one kilometer to ten meters by the optimum correction coefficient of thesmallest placement deviation.