节点文献

指令制导火箭弹最优控制弹道技术研究

Research on Optimal Control Trajectory Technology of Command Guided Rocket

部分文献由于文件较大,PDF全文下载时容易出现504错误,建议您优先选择CAJ下载或PDF分章下载。

【作者】 闫小龙

【导师】 陈国光

【作者基本信息】 中北大学, 兵器科学与技术, 2018, 博士

【摘要】 制导火箭弹因具有精度高、威力大、火力猛、射程远、成本低的诸多优点,受到世界各国的广泛重视。本文以鸭舵控制的旋转尾翼稳定式火箭弹为研究对象,围绕制导火箭弹研制过程中遇到的一些理论和技术问题,对飞行控制中的滚转姿态角测量和最优飞行弹道做了较为系统的研究,主要内容包括:综合考虑制导火箭弹高精度、抗干扰、低成本等方面的的设计要素,提出了基于无线电测量、磁罗盘、鸭式舵翼等技术的指令修正弹药姿态的技术和方案,主要讨论指令制导火箭弹的姿态测量与弹道优化技术。根据鸭式布局制导火箭弹的气动特性和弹道特征,建立了六自由度飞行动力学模型、滚转火箭弹的姿态控制系统模型、可操纵的火箭弹质点飞行运动方程。研制了相应的数值仿真分析软件。为了保障磁罗盘测量滚转角的整体精度,对磁罗盘测量滚转角的主要误差因素的作用进行了仿真研究。分别考虑了在无控弹道条件下由于射击诸元散布、发动机不一致性等因素造成的滚转角测量误差,以及在独立磁罗盘工作模式下的程控弹道中由于控制力的加入而造成的滚转角测量误差。最后分析了在无线电+磁罗盘测量闭环控制中制导模式下,磁罗盘测量弹体滚转角的精度范围,证明了在指令制导火箭弹中使用磁罗盘测角的优越性与可行性。建立了包含传感器误差、PCB制造偏差、结构安装偏差、弹体软硬磁干扰的线性磁测量模型;针对弹体捷联磁罗盘的校正问题,提出了基于椭圆拟合的两步快速可重构无迹卡尔曼滤波算法。该算法可实现地磁数字信号的实时处理,提取可用于舵机控制的弹体滚转姿态角和姿态角速度信息。数值仿真实验表明,该算法可在磁罗盘0.5ksps的采样频率下,在100ms内完成磁罗盘校正参数的精确辨识;半实物仿真实验表明,该滤波器在150ms内即可稳定输出弹体滚转姿态角,解算精度可达?0.8?。运用可操纵的制导火箭弹质点运动模型,以末端速度最大为目标函数,建立了中制导起控点弹道诸元、终点弹道诸元、控制量等约束条件的制导火箭弹中段弹道优化的变分模型。综合运用直接参数法和序列二次规划算法,给出了制导火箭弹的中制导最优弹道方案算法。通过仿真计算,给出了制导火箭弹在典型操纵能力下的弹道特征。中制导最优飞控弹道的规划设计工作要求在飞行弹道上的一个测控节拍内现场实时完成。火箭弹的飞行动力学模型存在强非线性、耦合性及多重约束条件,导致直接的轨迹优化算法的收敛速度过慢,解算时间太长。为了实时解算最优规划弹道,建立了规划弹道变分问题的有限弹道弧计算方法。该方法以有限弹道弧的形函数代替变分解法中的积分式,大幅度降低了算法的时间复杂度。仿真结果表明,对典型的弹道计算机,该算法可以满足规划弹道的实时解算要求。

【Abstract】 With high precision,strong power,long range,low cost and many other advantages,the guided rocket has attracted more and more attention around the world.In this paper,the canard controlled stabilized rocket with a rotating tail is studied,and some theoretical and technical problems for measuring roll attitude angle and optimizing flight trajectory during flight are discussed.The main contents include:Considering the design requirements of high precision,anti-jamming and low cost of the guided rocket,the technology and scheme for command correction of missile attitude based on radio measurement combining with geomagnetic compass and canard wings is proposed.The attitude measurement and the trajectory optimization of the command-guided rocket are mainly discussed.According to the aerodynamic and trajectory characteristics of the canard configuration guided rocket,the flight dynamics model with six degrees of freedom,the attitude control system model of the spinning rocket,the controllable particle motion equation of the rocket are established.The corresponding numerical simulation analysis software are developed.To ensure the overall accuracy of geomagnetic compass in measuring the roll angle,the main error factors in the measurement are analyzed by detailed simulation.The roll angle measurement error caused by the scatter of firing data and the motor inconsistency under the uncontrolled trajectory,and the control force addition in the programmable trajectory under the independent magnetic compass,are separately considered and studied.In addition,the precision range of the geomagnetic compass for measuring the roll angle of the missile under the closed-loop control mode of radio + geomagnetic compass is analyzed,which proves the superiority and feasibility of the geomagnetic compass in measuring the roll angle of the command-guided rocket.A linear magnetic measurement model for measuring sensor error,PCB manufacturing deviation,structural installation deviation,and soft and hard magnetic interference of themissile body is established.A two-step fast reconfigurable unscented Kalman filter algorithm based on ellipse fitting and UKF is proposed for the correction of missile strapdown geomagnetic compass.By this algorithm,the real-time processing of geomagnetic digital signal can be realized,in which the roll attitude angle and angular velocity can be extracted for controlling the steering engine.The numerical simulation results show that the algorithm can accurately identify the correction parameters of the geomagnetic compass in 100 ms at sampling frequency of 0.5 ksps.Moreover,the hardware-in-the-loop simulation is performed,and the results indicate that the roll attitude angle of the missile can be stably output in 150 ms with the calculation accuracy of ?0.8?.Based on the controllable particle motion model of the guided rocket,a variational model for the midcourse trajectory optimization of the guided rockets with minimum energy consumption for objective function is established.This model is restricted by the starting trajectory data,terminal trajectory data,and the steering range of the midcourse guidance.Combining the direct parametric method and sequential quadratic programming algorithm,the optimal trajectory scheme of guided rocket in midcourse guidence is obtained.The simulation calculation and comparative analysis are presented to get the trajectory characteristics of the guided rocket under typical control.The flight mathematical model of rocket existing strong nonlinearity,coupling and multiple constraints,which makes it difficult to select the initial value of direct trajectory optimization algorithm and slow the solve process.The planning and design of the optimal flight control trajectory of midcourse guidance requires real-time completion in a measurement and control rhythm on the flight trajectory.To solve the optimal planning trajectory,the calculation method of finite trajectory arc for planning trajectory variation is proposed.In this method,the finite trajectory arc function is used to replace the integral formula in the variational method,which greatly reduces the time complexity of the algorithm.The simulation results show that the algorithm can realize real-time calculation of planning trajectory with typical trajectory computers.

  • 【网络出版投稿人】 中北大学
  • 【网络出版年期】2019年 02期
节点文献中: 

本文链接的文献网络图示:

本文的引文网络
网页聊天
live chat
在线营销
live chat