共查询到19条相似文献,搜索用时 218 毫秒
1.
2.
3.
针对制导火箭落点速度的约束要求,提出了一种采用锥形运动控制导弹飞行速度的导引方法。该方法首先设计了满足速度约束的虚拟目标理想运动轨迹,将导弹减速控制问题转化为对虚拟目标的追踪导引问题,通过建立制导火箭与虚拟目标的相对运动模型,分析了弹目相对位置和相对速度的关系,推导了具有速度控制的导引律一般形式,并采用动态逆控制理论设计了锥形运动控制指令和导引参数。通过数字仿真对比了不同落角约束条件下导弹锥形运动的速度控制效果,结果表明该方法设计的导引律能够满足制导火箭速度约束要求,且制导精度高、控制效果好,为导弹锥形运动速度控制技术提供了参考。 相似文献
4.
液体火箭发动机推进弹道式导弹总体设计参数的全局最优化问题是亟待解决的计算问题。遗传算法具有全局搜索能力强、鲁棒性强、适于并行处理的特点,而Powell算法具有很好的求解局部最优解的能力。将两种方法进行有效改进后使之相结合,设计出并行全局最优化混合遗传算法。并以此为基础,建立了液体火箭发动机推进弹道式导弹总体优化设计模型。以液体火箭发动机推进弹道式导弹的起飞质量最小为目标,对液体推进剂弹道式导弹设计参数进行了优化设计。数值优化结果表明:该混合算法提高了搜索全局最优解的速度,优化精度高,且避免了初值敏感、病态梯度和局部收敛等问题,能够搜索到全局最优设计参数。 相似文献
5.
6.
本文将液体火箭发动机的动力循环过程分为开式循环和闭式循环,并通过分析的办法,建立了发动机动力循环的效率公式.在这些公式的基础上,分析了影响液体火箭发动机循环效率的因素.本文所得的结果,对液体火箭发动机设计计算有一定的参考价值. 相似文献
7.
8.
以液体火箭发动机整机试验、试车为目标,对发动机在研制过程中出现的各种故障进行了统计、归纳和分类。并按照目前液体火箭发动机试车测量参数的现状及不同的试车故障类型,制定了检测方法,也建立了故障分析模式。按照此种模式,在发动机未遭受到破坏之前及时停止试车,减少推进剂的浪费,使发动机的破坏尽可能地减到最少。 相似文献
9.
10.
11.
早期在对液体推进剂火箭发动机方案进行评价与选择时,仅以发动机本身的指标(如比冲、推重比等)作为方案比较的标准。这样没有考虑发动机子系统与运载器总系统的相互联系,得不到合理的评价结果。液体推进剂火箭发动机是航天运载器的一个子系统,采用运载器的性能指标评价发动机方案才能得到比较客观的结果。 本文推导了运载器的评价指标,给出了运载器的线性化质量方程,阐述了运载器设计参数的简化确定方法,由此提出了一个采用运载器评价发动机方案的方法。最后应用提出的方法对五个发动机方案进行了评价。 相似文献
12.
根据火箭发动机的推力公式及空气动力学有关理论,导出在小斜切角喷管内气体流动的特性、推力的变化规律,并给出喷管的设计方法. 相似文献
13.
《中国航空学报》2021,34(2):432-440
Reusable rocket engines are the core components of reusable launch vehicles, and have thus become a major focus of aerospace engineering research in recent years. In practice, subsystem design is based on the overall index allocation of an engine; therefore, a multidisciplinary optimization approach is necessary. In this study, design of a reusable methane/liquid oxygen (LOX/CH4) rocket engine with a gas generator cycle was investigated using multidisciplinary optimization. Two parameters were chosen as design variables: pressure and fuel mix ratio of the main combustion chamber. Optimization objectives were specific impulse, structural mass, and life cycle cost of the reusable rocket engine, and constraints were assigned to each discipline according to rocket design requirements. Then, an optimization model was developed, and optimal design parameters were acquired for the LOX/CH4 rocket engine. The proposed method is effective for designing the index allocation of reusable rocket engines and takes into account the multidisciplinary nature of complex systems. 相似文献
14.
火箭发动机整体结构及其部件的质量,质心位置和转动惯量的计算是火箭发动机及导弹的计算机辅助设计和优化问题中的重要组成部分。为此研制成功一种基于标准几何型体的位置,尺寸、方向和材料密度的交互式的火箭发动机结构质量特性计算软件MCM,给出了这些标准几何型体质量特性的精确计算公式。 相似文献
15.
以故障分析为目的,建立了一种大型泵压式液体火箭发动机的基本数学模型及实时数学模型,采用历史数据统计及数字仿真分析结合的方法,对发动机的故障模式及其效应进行了分析研究。提出了液体火箭发动机故障诊断系统的框架。为了对液体火箭发动机健康监控的算法及软件进行验证,以实时数学模型为基础,提供并建立了一个实时仿真验证系统。 相似文献
16.
V. A. Afanas’ev G. L. Degtyarev A. S. Meshchanov T. K. Sirazetdinov 《Russian Aeronautics (Iz VUZ)》2013,56(1):99-106
We obtained an on-off control law for two spacecraft (SC) rocket engines creating the angular acceleration in the opposite directions considering that SC is controlled by the rocket engines with an exponential increase of thrust at starting and decrease at shutdown. This control law provides the change-over of SC from the initial angular position to the final one in the scheduled time. A special trajectory consisting of eight typical segments is proposed to identify the angular acceleration and delay of its formation in the rocket engines. Both problems are solved by the method of analytical construction of composite trajectories for the angular turn from the typical segments on which the differential equations of angular motion have analytical solutions. The results are applicable for the development of promising space interceptors and design of onboard control systems. 相似文献
17.
讨论了带簇式推力室的固液火箭发动机多孔装药的设计方法,并针对某HTPB/LOX(或GOX)发动机给出了不同推力室数和不同孔数情形下某些参数随时间的变化,以及中一些参数的总体值或平均值,表明簇式推力室方案是可行的,研制适用于固液燃料配方也是必要的。 相似文献
18.
19.
High combustion temperatures and long operation durations require the use of cooling techniques in liquid propellant rocket engines (LPRE). For high-pressure and high-thrust rocket engines, regenerative cooling is the most preferred cooling method. Traditionally, approximately square cross sectional cooling channels have been used. However, recent studies have shown that by increasing the coolant channel height-to-width aspect ratio and changing the cross sectional area in non-critical regions for heat flux, the rocket combustion chamber gas-side wall temperature can be reduced significantly without an increase in the coolant pressure drop. In this study, the regenerative cooling of a liquid propellant rocket engine has been numerically simulated. The engine has been modeled to operate on a LOX/kerosene mixture at a chamber pressure of 60 bar with 300 kN thrust and kerosene is considered as the coolant. A numerical investigation was performed to determine the effect of different aspect ratio and number of cooling channels on gas-side wall and coolant temperatures and pressure drop in cooling channels. 相似文献