基于二次点火的飞行器横向转移入轨弹道设计

为了解决航天器发射过程因发射场地的地理位置约束造成的入轨夹角问题,基于火箭上面级二次点火,对航天器横向转移零速度偏差入轨弹道进行了设计。对二次点火和变射面横向转移弹道进行了研究,根据横向转移弹道特点,为简化控制程序,分时序对火箭二级非连续点火纵横向飞行程序进行了设计；并分析横向转移入轨弹道各项约束条件,建立弹道优化模型。根据弹道设计模型,以某两级液体燃料运载火箭为研究对象,对二次点火横向转移入轨弹道进行优化仿真。结果表明:入轨时刻航天器位置偏差为0.391m,速度偏差为1.277m/s,速度偏差比二次点火固定射面入轨弹道减少了737.844m/s,满足零速度偏差入轨精度要求。 

Trajectory design of launching plane transforming orbit injection of spacecraft based on engine twice ignition

In order to solve the velocity deviation of orbit injection question in space launch by the geographical location of launching site, the trajectory of zero velocity deviation orbit injection based on engine twice ignition and Launching plane transforming for the liquid launch vehicle was designed. According to the characteristics of the Launching plane transforming trajectory, the twice ignition and Launching plane transforming flight program model for the second stage was established in different sequences to simplify the control program of the engine, and the constraints of the trajectory were analysed to establish the trajectory optimization model. Then, the trajectory was simulated according to the optimization model. Results indicated that the orbit injection position deviation of the spacecraft was 0391m and the velocity deviation was 1277m/s, and the velocity deviation was 737844m/s less than the optimum result of the fixed launching plane orbit injection trajectory, meeting the requirement of the zero velocity deviation orbit injection.