Research On High Precision Trajectory Optimization And Guidance Algorithm For Launch Vehicle With Little Propellant Remaining

With the development of the aerospace industry,to be safe,reliable,fast-response and low-cost is the basic requirement for future launch vehicle.For safety and reliability,the launch vehicles need adaptive guidance technology with high adaptability and high robustness as a technical support.A high-precision fast trajectory optimization method can reduce set-up time,which makes it possible to be fast-response.In order to be low-cost,it is necessary to reduce the propellant remaining on the shutdown point,which improves the carrying capacity of the vehicle.Under this background,this paper srudys the trajectory optimization method and guidance method for launch vehicle with little propellant remaining.Modeling deviation factors and analysis for flight ability,carrying capacity and safety margin were carried out in the research.A dynamic model is established in both the geocentric inertial coordinates and the emission coordinates.The flight capability is described in two aspects: terminal index accuracy and carrying capacity.The carrying capacity enhancement,which is used as an index for guidance method performance,is defined as the increase of propellant remaining in shutdown point and the decrease of safety margin.The deviation factors affecting the flight state of the vehicle are modeled,and the influence of multiple deviation factors on the flight ability is quantitatively analyzed,which provides guidance for the subsequent research on the guidance algorithm.A high precision offline trajectory optimization method based on Legendre pseudospectral method is studied.Three analytical forms of flight procedures are given,and their optimizable parameters are listed.A serial composite optimization algorithm,composed with sequence quadratic programming(SQP)algorithm and genetic algorithm(GA),is used to optimize the parameters of three flight procedures,which takes various path constraints into consideration,such as dynamic head constraint and normal overload constraint.Based on Legendre pseudospectral method,a full ballistic optimization model is established,and the continuous optimal control problem is transformed into a parameter optimization problem,which is solved with the sequence quadratic programming algorithm.An adaptive guidance method based on analytic solution is studied.At first,the theory of an iterative guidance method is analylized.This iterative guidance method is improved for specific cases and a double-loop iterative guidance model is constructed.Based on the simplified dynamic equations in the local coordinates,with the derivative value of the pitch angle’s trigonometric function as the control variable,the analytical solution that satisfies the Pontryagin maximum principle is deduced.And based on this analytical solution,online adaptive guidance method was constructed.This method guarantees continuity of the attitude angle between the flight stages and takes the terminal angle of attack constraint into account.The numerical simulation results show that compared to the iterative guidance method,this kind of adaptive guidance method can guarantee more accurate terminal state for launch vehicle,and has stronger anti-interference capability,which improves the launch vehicle’s carrying capacity.The application of aerodynamic parameters online identification based on Extended Kalman Filter in guidance calculation is studied.Based on the simplified dynamic equations in the velocity coordinates,an Extended Kalman Filter for aerodynamic parameters identification is constructed,in which the identification variable is the aerodynamic coefficient deviation.The perturbation equation considering the aerodynamic coefficient deviation is derived,and the aerodynamic parameter identification result is applied to the guidance calculation.The simulation shows that the filter converges rapidly and has fairly high precision.The aerodynamic parameter identification can effectively reduce the terminal state deviation,and improve the convergence of the trajectory,which improves the guidance method performance.The work of this paper can provide support for the trajectory optimization and on-line guidance for launch vehicle in the case of low propellant margin.It can provide reference for the application of aerodynamic parameter identification in guidance and trajectory optimization and guidance in the atmosphere.