Earth-to-Moon missions with low thrust-to-weight ratios present unique challenges for exoatmospheric guidance, and the existing algorithms are ineffective for the unprecedentedly long burn arcs and high orbital eccentricities. To address these challenges, a Long Burn Arc Powered Explicit Guidance (LBA-PEG) algorithm is developed and compared with the existing algorithms. In the proposed LBA-PEG algorithm, a fully numerical thrust prediction method is developed to accurately predict the highly nonlinear thrust effects over long burn arcs. Moreover, a real-time Newton correction method is proposed to correct the orbit injection point, remedying the position-velocity coupling induced by high orbital eccentricities. The comparison between the proposed algorithm and the existing algorithm shows that the proposed algorithm surpasses the existing ones by significantly enhancing fuel efficiency and improving tolerance to thrust decrease. The proposed LBA-PEG algorithm can adapt to a 65% thrust decrease, which is 12%-22% larger than that of the existing algorithms, and it can still reliably converge and complete the guidance mission even when the length of the burn arc exceeds 90 degrees. The proposed LBA-PEG highlights the algorithm's adaptability for long burn arc missions, especially in critical scenarios such as manned Earth-to-Moon missions. (c) 2025 Published by Elsevier Ltd on behalf of Chinese Society of Aeronautics and Astronautics. This is an open access article under the CC BY-NC-ND license (http://creativecommons.org/licenses/by-nc-nd/4.0/).

To cope with the potential thrust-drop malfunction of the manned lunar exploration launch vehicle, a parking orbit replanning method is proposed based on the evaluation of the residual carrying capacity of the launcher. By analyzing the constraints of the whole flight profile and the characteristics of the Earth–Moon transfer orbit systematically, the offline trajectory planning algorithm exhibits improved convergence performance, and thus can be used to evaluate the residual carrying capacity as thrust-drop happens. For the situations that the launcher is not capable of injecting the transfer orbit, a sequential orbit/trajectory replanning method is designed to guarantee the safety of the astronauts. On the premise of ensuring the altitude of the parking orbit, the elements about the orbital plane are further optimized to provide a favorable initial state for the subsequent rescue action. Meanwhile, the nonlinearity of the terminal constraints is alleviated by the injection point forecast, and the initial reference is generated by convex optimization method with well convergence; thus, the solving efficiency of the sequential replanning algorithm can be improved. Simulation results show the proposed method can generate the trajectory that transport the spacecraft to the optimal parking orbit under the thrust-drop malfunction situation. This is an English translated version of “Residual Carrying Capacity Evaluation and Parking Orbit Re-planning for Lunar Exploration Launch Vehicle”.

Aiming at the load relief method of the new generation launch vehicle during its boost phase when flying through windy areas. The multi-loop load relief method based on overload feedback loop and extended state observer(ESO) is studied theoretically. The control parameters are tuned by particle swarm optimization algorithm(PSO), and an improved algorithm with particle reset strategy is proposed to improve the population diversity and avoid the population falling into local optimum. The simulation evaluation results show that the active load relief method based on the improved particle swarm optimization algorithm has better adaptability to high-altitude wind disturbance and can further improve the load relief capability of the launch vehicle.

Aiming at the problem that the shear wind interference makes the aerodynamic load larger when the new generation launch vehicle passes through the windy area, a multi-loop self-adaptive active load relief control method is proposed. Based on the overload feedback load relief method, the observer compensation loop is introduced, and the influence of the multi-loop joint load relief control method on the aerodynamic load and the control load is derived. An adaptive gain real-time scheduling algorithm based on aerodynamic estimation is designed. The simulation evaluation results show that the multi-loop adaptive load relief control method has better adaptability to shear wind interference than the overload feedback load relief control method, which can further improve the load relief ability.

运载火箭通常采用捷联惯组测量视加速度和角速度,通过导航系统计算火箭的速度、位置和姿态,用于制导和控制.为了确保重型运载火箭控制系统导航信息的高可靠性,提出了一种基于双十表光学捷联惯组的主从冗余信息管理方法及信息重构技术.该方法可对双五表陀螺仪与双五表加速度计配置下的主从惯组冗余信息进行故障判别和隔离,利用有效测量信息进行冗余信息重构,再依据寻优的设计原则完成冗余信息的选用.对于飞行过程中惯组的同类单表发生三度以下故障的情况,利用该方法可通过系统级冗余实现火箭控制系统高精度可靠工作,确保完成发射任务.

AbstractThe purpose of the guidance control is to release a payload into a prescribed target orbit (PTO) accurately. The parameters that determine an orbit are called orbital elements (OEs), which include the semi-major axis a, the eccentricity e, the argument of perigee $$\omega $$ ω , the inclination angle i, and the longitude of ascending node (LAN) or the right ascension of ascending node (RAAN) $$\Omega $$ Ω , where a and e can be converted to the perigee height $$h_p$$ h p and the apogee height $$h_a$$ h a . Thus, the guidance mission of a launcher is a typical optimal control problem with multi-terminal constraints, which requires complex iterative calculations. Considering various constraints in practical applications, such as the accuracy of inertial navigation systems and the performances of embedded computing devices (speed and storage capacity), guidance methods need to balance the mission requirements, hardware resources, and algorithm complexity. A variety of guidance methods has been developed with distinct era characteristics.

This paper presents a study of the autonomous guidance reconfiguration of a launch vehicle, which improves the adaptability of the launch vehicle after the thrust drop fault with the online replanning of the transfer orbit and the shutdown and startup conditions of the guidance method. For a multi-satellite launch mission with a coast phase, we analyze the guidance constraints in the first powered phase, coast phase, and second powered phase, and we propose an online multi-grade optimization to separate some satellites during the coast phase so that the remaining satellites can still enter the original target orbit. We then use the adaptive collocation method to solve the replanning problem, study the sparsity of the discrete matrix, and develop an initial guess generation method for an engine out fault to improve the computational efficiency. The numerical simulation shows that the proposed method, which can partially complete the launch mission with severe faults by sacrificing some satellites, further enhances the reliability of the launch vehicle.

For missions that involve launching satellites into sun-synchronous orbits, the load-carrying capacity can be improved by regulating the time distribution of different flight phases if a transfer orbit is planned. However, this will shorten the last burn time of the engines. This paper introduces the technologies to ensure a reliable engine restart after coasting when only a reduced level of propellants is left. The influence mechanism of different flight profiles on the carrying capacity is analyzed first. Then a joint simulation method integrating the attitude control and fluid dynamics is proposed that can accurately analyze the sloshing features of the cryogenic propellants and their influences under microgravity conditions. Finally, a guidance method is studied that eliminates the attitude deviations at the shutdown time of the first burn to suppress additional propellant sloshing during the follow-up coasting phase. The high-fidelity simulation system reveals the feasibility of a low level scenario, and with the action of the updated guidance method, the pressure drop can be controlled as needed and the sloshing can be well restrained, matching the existing pressurization ability and requirements for re-ignitions of engines. These technologies are applied in the successful launch of the LM-8/Y2 mission, in which the performance improvement is verified.

This paper describes the study of a powered-coast-powered guidance reconfiguration (PCPGR) method used to solve the autonomous rescue problem for the mission profile of a launch vehicle with a coasting phase when the thrust drop fault occurs in the first powered phase (FPP) of the final stage. We first described the constraints of the final stage and the construction of the PCP guidance problem. Then we evaluated the adaptability of the guidance reconfiguration (GR) offline with numerical optimization by adjusting the constraints of the FPP, the coast phase, and the second powered phase. To determine the fault state set where the rocket can enter the prescribed target orbit through the GR initiate, we proposed a Newton method-based rapid replanning method of the transfer orbit that transforms the complex multi-flight phase trajectory planning problem into a feasible transfer orbit search problem to produce a fast solution onboard. Combined with the adaptive adjustment of the coasting time and the iterative guidance mode, we realized the autonomous online rescue of the payload. The simulation results showed that the proposed method achieved a reliable and rapid solution and improved a launch vehicle’s adaptability to a thrust drop fault.

Facing diversified launch needs, it is uneconomical to redevelop new rockets from scratch. The concept of combining legacy modules into a new launcher has become attractive. However, these mature modules may not adapt to new flight profiles due to differences in factors such as the propulsion system configuration, payloads, and trajectories. Thus, structural load relief measures play a key role in module integration. A comprehensive load relief strategy applied in the Long March 8 (LM-8) rocket is introduced, including the inflight load reduction by engine throttling, wind biasing trajectory, and onboard load relief control techniques. A unified analysis process for the elastic loads caused by gusts and fluctuating pressures was proposed, and an integral optimization problem for simultaneous planning with complicated constraints (such as the maximum dynamic pressure, wind biasing, and throttling level and time) is discussed. A real-time load relief technology based on an extended state observer, which predicts the angular acceleration caused by the wind, is proposed. Its efficiency, adaptability, influence on the stability and control accuracy, and application in the LM-8 maiden flight are studied. The proposed scheme expands the launch probability of LM-8 and provides a systematic solution for the load relief design and control of in-service rockets.

为提高运载火箭在飞行中非致命故障和大偏差条件下的自适应能力,对运载火箭非致命故障下弹道规划制导和自适应重构控制国内外技术现状进行了分析,提出了火箭自主飞行重构方案,重点研究了轨迹规划方法制导化的弹道规划制导、控制参数自适应切换策略和控制指令重分配方法,开展了自主飞行重构仿真分析,完成了工程应用研究.仿真结果表明:自主飞行重构方法在运载火箭上可行,能够提升火箭应对不确定性和非致命故障的适应性.

In recent years,Chinese Long March(LM)launchers have experienced several launch failures,most of which occurred in their propulsion systems,and this paper studies Autonomous Mission Reconstruction(AMRC)technology to alleviate losses due to these failures.The status of the techniques related to AMRC,including trajectory and mission planning,guidance methods,and fault tolerant technologies,are reviewed,and their features are compared,which reflect the challenges faced by AMRC technology.After a brief introduction about the failure modes of engi-nes that can occur during flight,and the fundamentals of trajectory planning and joint optimization of the target orbit and flight path,an AMRC algorithm is proposed for geostationary transfer orbit launch missions.The algorithm evaluates the residual performance onboard,and plans new objec-tives and corresponding flight path by iterative guidance mode or segmented state triggered opti-mization methods in real-time.Three failure scenarios that have occurred during previous LM missions are simulated to check the robustness of the algorithm:imminent explosion risk of the boosters'engines,thrust drop during the first stage of flight,and being unable to start the engine during the second stage.The payloads would fall from space according to the current design under these conditions,but they were saved with the AMRC algorithm in the simulations,which allowed the rocket to get into the target orbit as intended or the payloads were deployed in other orbits with-out crashing.Although spaceflight can be very unforgiving,the AMRC algorithm has the potential to avoid the total loss of a launch mission when faced with these kinds of typical failures.

This paper reviews the development of reliability design technology as well as challenges encountered thus far regarding control systems for long march launch vehicles (LMLVs). Along with the high frequency of space launches, the target of “increasing the success rate by one percentage point” has been proposed by the China Academy of Launch Vehicle Technology for LMLVs. In this context, the features of the representative launchers in the reliability design of control systems are reviewed first, especially from an industrial perspective. Thereafter, the techniques applied by LMLV control systems are presented using a five-level hierarchical architecture: fundamentals, self-repair, multiple module redundancy, fault-tolerant control, and fault-tolerant guidance. This is followed by a detailed discussion on the reliability design in LMLV navigation, guidance, control, and avionics subsystems. Finally, the challenges are addressed, which encompass the cost-oriented reliability design, the fault-tolerant control to suppress out-of-tolerance behaviors, and the fault-tolerant guidance to replan the mission autonomously under critical faults. The techniques under development for LMLVs in these areas are also summarized.

This paper proposes a guidance method for the powered soft landing of a launcher with non-cluster configured engines, for which it is difficult to maintain a low thrust-to-weight ratio (TWR) preferable for rocket recovery. For rockets in service with one or two engines rather than a cluster of engines in the boosters, it is a challenge to deeply throttle the engines to a low level, especially lower than 30%, and if one of the two engines is shut down to reduce the thrust, the attitude control is also difficult due to thrust asymmetry. The theoretical analysis and simulations show that the higher the TWR is, the smaller the feasible region for a soft landing becomes. The analysis also reveals that the bang–bang solutions of fuel-optimal methods are vulnerable to uncertainties and disturbances. A thrust regulation rate constraint is proposed or convexified to reconstruct the landing problem under high TWRs. An initial guess based on mean thrust and relaxing the initial velocity constraint is proposed, making both the adaptive collocation method and the successive convex programming applicable for real-time landing planning. Simulations under TWRs larger than four show that the proposed methods ensure a safe landing and are more adaptive than fuel-optimal solutions when facing disturbances and uncertainties.

This paper studies the online trajectory planning method of the rocket vertical landing under the condition of high thrust-to-weight ratio (HTWR). This is common for most of the existing rockets which are configured with high-thrust engines but only few are installed. The HTWR leads to a substantial reduction in the physical feasible region of the optimal solution. Moreover, the traditional fuel optimal strategy has the characteristics of bang-bang control, which makes the rocket weak adaptability to disturbance or model deviation in the state of maximum or minimum thrust. Thus the online planning could be unsolvable or violate the feasible boundary. This paper first describes the equations of motion and constraints during the landing process, and then the influence of the HTWR is analyzed from the perspectives of physical feasible region and deviation adaptability. In this paper, an optimal objective function with larger feasible region is proposed, so that the online planning trajectory under the condition of HTWR can be far from the boundary of the feasible region. The nonlinearity of the problem is reduced by the method of compensation model, and a second-order cone programming problem with free terminal time suitable for the primal dual interior point method is constructed. The simulation results show that the landing trajectory planned by the proposed algorithm is adaptive to the model uncertainty and disturbances, and is more robust to achieving a safe landing with a HTWR.

本文对重复使用运载器制导与控制技术进行综述.随着航天技术的发展,对航天运载器重复使用的需求也日益剧增,具备可复用的天地往返运输能力也一直是航天工业追求的重要目标之一,而制导与控制将发挥重要的作用.首先回顾了全球范围内重复使用运载器的研究进展,随后从不同的维度对其发展途径进行分类和分析,并从垂直起飞垂直着陆(VTVL)、垂直起飞水平着陆(VTHL)、水平起飞水平着陆(HTHL)等3个方面对制导与控制的需求进行了梳理.针对不同的起降模式,详细构建了完整的制导与控制模型、约束与目标函数,从而对比在不同场景下制导与控制的特点和挑战.在此基础上,对在VTVL、VTHL、HTHL 3种工作方式下制导与控制理论研究与工程实践中所取得的研究成果进行分析,并对各种方法的特点进行了论述和比对.最后对本领域当前亟待突破的技术难点和发展趋势进行了讨论,并对推动重复使用运载器应用的重点研究方向进行了归纳和展望.

子级垂直回收的一项关键技术是制导、导航与控制(Guidance Navigation and Control,GNC)技术.瞄准未来的组合体回收模式,结合现阶段发动机、着陆支架等环节的性能和特点,全面分析了返回过程中最关键的垂直着陆段对控制系统的需求和约束条件,研究了一种针对垂直着陆段的GNC技术,包含高精度导航算法、适应大推重比运载器快速下降模式着陆的凸优化在线轨迹规划与制导算法、适应大静不稳和晃动极零结构的参数优化设计以及基于自抗扰的高精度姿态控制算法.仿真结果表明,提出的技术可以满足设计需求.

本文综述了运载火箭上升段自主制导方法(autonomous guidance method,AGM)的发展.AGM是指一类不依赖离线规划的参考轨迹,而是实时规划满足后续飞行复杂过程约束和终端条件的飞行轨迹,从而动态计算当前制导指令的方法.该方法结合了轨迹规划和传统制导方法的优点,具有在线、动态、全局和滚动的特点.分析了不同运动体在制导控制上面临的主要挑战,提出了长征运载火箭上升段自主制导方法的两大技术体系.一是针对确定目标轨道的制导方法,也称迭代制导(iterative guidance mode,IGM);介绍了迭代制导的基本理论、基础型及增强型算法,及其在型号中应用的飞行结果.二是目标轨道与飞行轨迹的联合优化方法,主要应对推力下降的典型故障;介绍了基于目标函数触发(state triggered indices,STI)的在线优化方法,以及以STI和IGM为基础进行自主任务重构的算法.论文还提出了未来的研究方向,包括多级分层优化,以及基于自主制导方法的新型闭环控制系统架构,以进一步增强上升段飞行的自主性和故障适应性.

介绍了新一代长征运载火箭(LMLVs)的型谱,并从四个方面对运载火箭控制系统的发展进行了综述.制导技术从开环制导发展到迭代制导,并针对大推力直接入轨和终端姿态约束要求,进一步发展了迭代制导算法,入轨精度大幅提升;姿态控制仍以PID技术为基础,采用空间模态和等效摆角的建模方法解决助推飞行段多个舱段发动机联合摇摆问题,结合自抗扰技术(ARDC)进行主动减载控制;自载人航天工程起开展系统性的可靠性设计研究,逐渐形成了以设备冗余、算法容错和系统在线重构等为特点的技术体系,促进了长征火箭控制系统可靠性的整体提升;电子系统从分立的集中式体系架构,发展为集成化的分布式数字控制系统.针对当前飞行控制技术的研究热点,本文最后总结了长征运载火箭在这方面的最新实践与发展趋势.

This paper summarizes the autonomous guidance methods (AGMs) for pinpoint soft landing on celestial surfaces. Wefirst review the development of powered descent guidance methods, focusing on their contributions for dealing with constraints and enhancing computational efficiency. With the increasing demand for reusable launchers and more scientific returns from space exploration, pinpoint soft landing has become a basic requirement. Unlike the kilometer-level precision for previous activities, the position accuracy of future planetary landers is within tens of meters of a target respecting all constraints of velocity and attitude, which is a very difficult task and arouses renewed interest in AGMs. This paper states the generalized three- and six-degree-of-freedom optimization problems in the powered descent phase and compares the features of three typical scenarios, i.e., the lunar, Mars, and Earth landing. On this basis, the paper details the characteristics and adaptability of AGMs by comparing aspects of analytical guidance methods, numerical optimization algorithms, and learning-based methods, and discusses the convexification treatment and solution strategies for non-convex problems. Three key issues related to AGM application, including physical feasibility, model accuracy, and real-time performance, are presented afterward for discussion. Many space organizations, such as those in the United States, China, France, Germany, and Japan, have also developed free-flying demonstrators to carry out related research. The guidance methods which have been tested on these demonstrators are briefly introduced at the end of the paper.