运载火箭地月转移入轨混合制导方法

郭康; 李惠峰; 包为民

doi:10.13700/j.bh.1001-5965.2025.0353

运载火箭地月转移入轨混合制导方法

doi: 10.13700/j.bh.1001-5965.2025.0353

郭康^{1, 2},
李惠峰^{1, 2, ,},
包为民³

1.
北京航空航天大学宇航学院，北京 100191
2.
北京航空航天大学天地往返高效运输技术全国重点实验室，北京 100191
3.
中国航天科技集团有限公司，北京 100048

基金项目:

国家自然科学基金(92471204)

详细信息

通讯作者:
E-mail：lihuifeng@buaa.edu.cn

中图分类号: V448.13
计量
- 文章访问数: 230
- HTML全文浏览量: 112
- PDF下载量: 8
- 被引次数: 0
出版历程
- 收稿日期: 2025-06-06
- 录用日期: 2025-12-05
- 网络出版日期: 2025-12-17
- 整期出版日期: 2026-10-31

Hybrid guidance method for launch vehicle trans-lunar injection

GUO Kang^{1, 2},
LI Huifeng^{1, 2
, ,},
BAO Weimin³

1.
School of Astronautics，Beihang University，Beijing 100191，China
2.
State Key Laboratory of High-Efficiency Reusable Aerospace Transportation Technology，Beihang University，Beijing 100191，China
3.
China Aerospace Science and Technology Corporation，Beijing 100048，China

Funds:

National Natural Science Foundation of China (92471204)

More Information

Corresponding author: E-mail：lihuifeng@buaa.edu.cn

摘要

摘要:
针对小推重比地月转移入轨任务中长飞行时间和强不确定性对运载火箭闭环制导方法带来的挑战，提出一种混合使用二阶锥规划和邻近最优的制导方法。建立考虑多源不确定性的地月转移入轨问题模型，并提出一种基于实时轨迹规划的混合制导策略。实现一种基于高斯伪谱离散与二阶锥规划的实时轨迹规划方法，综合应用协态映射定理和二阶锥规划对偶输出，能够高精度地估计协态量，为邻近最优方法提供必要基础。提出一种基于邻近最优的轨迹校正方法，在不确定性导致火箭偏离轨迹时，能够为二阶锥规划提供高质量初始参考。开展数值实验，在一个地月转移入轨任务中，与现有制导方法对比，所提制导方法对地月转移入轨任务具有更强的适应性和最优性，混合制导方法的平均计算时间降低30.67%，每制导周期的平均算时间为75.353 ms，为地月转移入轨提供了一条新且可行的技术路线。
- 运载火箭制导 /
- 地月转移入轨 /
- 实时轨迹规划 /
- 二阶锥规划 /
- 邻近最优
Abstract:
This paper addresses the critical guidance problem for launch vehicles performing trans-lunar injection missions with low thrust-to-weight ratios, where the extended flight duration and substantial multi-source uncertainties present formidable challenges to conventional closed-loop guidance methods. A novel hybrid guidance approach integrating second-order cone programming with neighboring optimal control is developed to overcome these challenges. The proposed methodology begins with establishing a comprehensive trans-lunar injection problem model that explicitly accounts for multi-source uncertainties. A hybrid guidance strategy based on real-time trajectory planning is formulated to handle these uncertainties effectively. The main innovation is the application of a real-time trajectory planning technique that uses second-order cone programming in conjunction with Gauss pseudospectral discretization. By synergistically applying the costate mapping theorem and exploiting the dual output of second-order cone programming, the method achieves high-precision costate estimation, which provides the essential foundation for the neighboring optimal control component. A trajectory correction method based on neighboring optimal control is developed to address situations where uncertainties cause significant trajectory deviations. This method generates high-quality initial references for the second-order cone programming solver, substantially improving convergence characteristics and computational efficiency. Extensive numerical experiments demonstrate the superiority of the proposed guidance method in trans-lunar injection missions. The computational performance shows remarkable improvement, with the average computation time reduced to 75.353 milliseconds per guidance cycle, representing a 30.67% reduction. In comparison to current methods, this hybrid guidance method’s successful demonstration offers a novel and feasible technical pathway for trans-lunar injection missions.
- launch vehicle guidance /
- trans-lunar injection /
- real-time trajectory planning /
- second-order cone programming /
- neighboring optimal control

HTML全文

图 1 基于实时轨迹规划混合制导方法的流程

Figure 1. Flowchart of a hybrid guidance method based on real-time path planning

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图 2 混合制导方法示意图

Figure 2. Schematic diagram of a hybrid guidance method

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图 3 高度与速度曲线

Figure 3. Altitude-versus-speed curve

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图 4 控制量曲线

Figure 4. Control variable curve

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图 5 与纯二阶锥规划方法的计算时间对比

Figure 5. Comparison of computation times with pure second-order conic programming methods

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图 6 推力进一步下降时的高度与速度曲线

Figure 6. Altitude-versus-speed curve as thrust decreases further

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图 7 推力进一步下降时的控制量曲线

Figure 7. Control variable curve as thrust decreases further

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表 1 不同动力故障下的入轨精度

Table 1. Orbital insertion accuracy under different cases

工况	偏心率偏差	轨道倾角偏差/(°)	升交点赤经偏差/(°)	近地点幅角偏差/(°)
下降40%	0.000200	0.000006	0.000015	−0.004397
下降45%	0.000270	0.000005	0.000016	−0.004574
下降50%	0.000373	0.000004	0.000017	−0.004923

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