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31 March 2026, Volume 46 Issue 2 Previous Issue   

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Research on system design and key technologies for crewed lunar landing and ascent WANG Xiang, ZHU Enyong, LIU Yang, HOU Zhendong 2026, 46 (2):  1-11.  doi: 10.16708/j.cnki.1000-758X.2026.0019 Abstract ( 437 )   PDF (8772KB) ( 695 )   Save Lunar landing and ascent represents one of the riskiest and most challenging phases in crewed lunar exploration missions, exerting a direct impact on mission success and astronaut safety. Mastering and breaking through the key technologies for crewed lunar landing and ascent formed the core of developing a safe and reliable crewed lunar lander, and further drove the overall advancement of China’s crewed spaceflight technology. First, mission requirements for crewed lunar landing and ascent were analyzed, followed by the proposal of a system scheme for China's crewed lunar lander. Based on this scheme, an in-depth analysis was conducted on the key landing and ascent technologies specific to typical crewed mission demands, including emergency rescue design, ultra-soft landing design, lunar surface stable ascent, and autonomous rapid circumlunar rendezvous. Corresponding technical approaches and solutions were put forward accordingly. A circumlunar orbit design method optimizing fuel consumption for both nominal and emergency ascent, as well as an integrated planning method for emergency ascent and rendezvous-docking adapting to multi-constraint and variable-target scenarios were proposed, effectively enabling the crew to return safely at any stage during the landing and ascent process. A multidisciplinary optimization method for descent velocity was developed, taking into account plume heating, landing impact and landing stability, which provided a design reference for achieving ultra-soft lunar landing. A thrust coordinated control method for parallel orbit control engines and a rapid engine fault diagnosis method were established, enhancing the lander’s control capability and safety during lunar ascent. Additionally, a scheme design method for autonomous rapid circumlunar rendezvous under complex scenarios and multi-constraint conditions was presented, addressing the challenges in rendezvous scheme design under resource and time limitations. This analysis of key technologies for lunar landing and ascent provides a reference for the design of crewed lunar landers, and also helps deepen the understanding of the development difficulties and risks associated with such landers. Related Articles | Metrics

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Integrated design method for load of spacecraft and launch vehicle ZHANG Zhi, WANG Lei, LIU Hui, HU Xiaojun, WANG Ziyang 2026, 46 (2):  12-20.  doi: 10.16708/j.cnki.1000-758X.2026.0020 Abstract ( 148 )   PDF (9856KB) ( 238 )   Save To address issues such as aerodynamic coordination design, challenges in matching fundamental frequencies, conservative load design, and over-testing in ground experiments exposed by the traditional independent design mode of spacecraft and launch vehicles when dealing with large-scale spacecraft, an integrated spacecraft-launch vehicle collaborative design method is proposed. For joint optimization of aerodynamic configuration: by comparing and analyzing the pulsating pressure distribution of the self-escape and launch escape tower configurations, the optimal solution is determined. For collaborative dynamic characteristics design: a coupled spacecraft-launch vehicle dynamics model is established. To address the ultra-low fundamental frequency of the spacecraft (2.8Hz) and the dense elastic frequencies of the launch vehicle, a combination of structural stiffness enhancement and control system optimization method is adopted. For engineering extrapolation of external force functions during liftoff: based on jet flow simulations, external load identification, and historical flight data, an extrapolation method from noise to external forces is developed.For refined load design: a load analysis method integrating surface aerodynamic distribution, static overload, and dynamic acceleration is proposed. Dynamic response analysis is conducted based on the mechanical conditions of module interfaces, replacing the traditional centroid-equivalent overload method. The application of this method has yielded clear results. In terms of aerodynamics, the peak pulsating pressure at key locations of the launch escape tower configuration is reduced by approximately 50% compared with the self-escape configuration. In terms of dynamic characteristics, after optimization, the first and second elastic frequencies of the launch vehicle are effectively separated. The amplitude of the first three elastic frequency domains in the attitude control system is reduced by up to 4.10dB, significantly improving stability. In terms of loads and testing, the test load spectrum generated based on module interface conditions is reduced by an average of approximately 20% compared with traditional fixed-base conditions for the complete vehicle, effectively mitigating ground over-testing issues. The integrated spacecraft-launch vehicle design method proposed in this paper achieves multidisciplinary optimization across aerodynamic configuration, structural dynamics, attitude control design, and load environment, breaking through the limitations of traditional interface-segmented design. Its core value lies in providing systematic solutions for addressing ultra-low fundamental frequency compatibility and complex load prediction, while offering an effective approach for spacecraft weight reduction and reliability improvement through margin sharing and refined design. Related Articles | Metrics

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Research on human lunar exploration mission replanning under fault conditions ZHANG Hailian 2026, 46 (2):  21-35.  doi: 10.16708/j.cnki.1000-758X.2026.0021 Abstract ( 164 )   PDF (3637KB) ( 220 )   Save China's manned lunar exploration (MLE) project employs a circumlunar rendezvous and docking flight plan, which involves multiple spacecraft and flight phases. Given the system's complexity, various risk factors, and high mission costs, it is essential to consider mission replanning under various fault scenarios to ensure astronaut safety and maximize mission benefits. Firstly, the mission replanning problem is defined, categorizing it into four types: mission reconstruction, mission adjustment, mission degradation, and mission abortion. Subsequently, the studies on replanning methods across different flight phases of the MLE mission are systematically reviewed, from the launch phase to the lunar return phase. Then, four key technologies that still require breakthroughs to address the replanning problem are concluded, including fault diagnosis, detection and prediction, multi-objective optimization and degrading objective decision, integrated mission planning, and highly robust and real-time mission replanning technologies. Finally, the characteristics and challenges of the replanning problem are summarized and a forward-looking perspective on future developments is offered. Related Articles | Metrics

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A reachable domain calculation method of contingency return trajectories based on interval number PENG Qibo, LU Lin, ZHANG Hailian 2026, 46 (2):  36-47.  doi: 10.16708/j.cnki.1000-758X.2026.0022 Abstract ( 81 )   PDF (8366KB) ( 124 )   Save A reachable domain calculation method based on interval number is proposed for the contingency return trajectories during the circumlunar flight phase. Firstly, the problem of a contingency return trajectory which adopts the non-coplanar maneuver during the circumlunar flight phase is described. The mathematics model of the reachable domain of the contingency return trajectories is established. Secondly, based on the design of the contingency return trajectory, a reachable domain calculation method of contingency return trajectories is proposed by introducing an improved interval branch and bound algorithm. Thirdly, the comparison with the traditional reachable domain calculation method based on brute force is performed in the simulations. The comparison result indicates that this method can reduce the calculation time by at least 47.4% and by at most 96.5%, and verifies the accuracy and efficiency of the proposed method. Finally, extensive simulation calculations are performed to analyze the parameter characteristic of reachable domain by using this method. The research conclusions can provide reference for the selection and decision of contingency return scheme in the future lunar exploration. Related Articles | Metrics

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Mechanical analysis of takeoff safety of 0-0 launch abort test for Mengzhou manned spacecraft LI Xisheng, WU Wenrui, LIANG Fuwen, HAO Baoxin, YAN Jun, TIAN Lin, MA Xiaobing 2026, 46 (2):  48-56.  doi: 10.16708/j.cnki.1000-758X.2026.0023 Abstract ( 104 )   PDF (2974KB) ( 225 )   Save At the takeoff time of the 0-0 launch abort test for Mengzhou manned spacecraft, multiple key processes are parallel. The potential fault modes of the key processes are coupled with each other, resulting in complex fault mechanisms and unclear mechanical responses. By analyzing the mechanical behavior of each fault mode at the takeoff time of the test, the takeoff safety of the 0-0 abort test is evaluated.The research is mainly conducted through theoretical analysis and numerical simulation. According to the force conditions of the escape system under different fault modes, the theoretical formula for the local maximum load of the escape system is derived. At the same time, finite element analysis and dynamic analysis are combined to confirm the theoretical results. Based on the mechanical analysis results, whether the escape system under different fault modes has the structual damage, whether it owns escape capabillity, as well as the impact on the trajectory after escape are given. Finally, the takeoff safety is evaluated through margin analysis.Through mechanical analysis, there are four fault modes, the safety margin of which is greater than 0.7 during the takeoff of the 0-0 launch abort test. Only one fault mode has no safety margin and it is classified as a medium level risk.Through a systematic study on the takeoff safety of the 0-0 launch abort test of Mengzhou manned spacecraft, a set of mechanical analysis method for the escape system is proposed: a theoretical model for the axial load of the cabin pyrotechnic lock is established; the methods of finite element analysis and dynamic analysis for each fault mode are proposed;the key mechanical characteristics that affect the safety of the escape system are identified. The research method in this paper can be extended to the takeoff safety analysis of the escape tests or launch missions for similar spacecraft. Related Articles | Metrics

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Integrative and autonomous control design for rendezvous and docking of Mengzhou spacecraft LIU Zhiyong, CAI Biao, FENG Ye, ZHANG Yi, SU Yan 2026, 46 (2):  57-70.  doi: 10.16708/j.cnki.1000-758X.2026.0024 Abstract ( 120 )   PDF (10183KB) ( 186 )   Save Mengzhou spacecraft includes an Earth version and a lunar version, which respectively perform rendezvous and docking with the Chinese space station in earth orbit and the Lanyue spacecraft in lunar orbit. During the design process of Mengzhou spacecraft, two prominent challenges emerged: deep integration of the autonomous high fault-tolerance GNC systems for rendezvous and docking with the Chinese space station and the Lanyue spacecraft, autonomous guidance, navigation and control during rendezvous and docking process. In this study, the specific situation of two problems is described, the problems' solutions are given, and the solutions' effects are verified by ground testing. The results indicate that the rendezvous and docking of the Mengzhou Earth spacecraft and the Mengzhou lunar spacecraft adopt an autonomous fault-tolerant system based on TTE and 1553B hot backup communication and four-model redundant computer, achieving the goal of continuous fault-tolerant operation under one fault, continuous safe operation under dual faults. Through autonomous navigation with multiple sensors and multi-stage autonomous coordination guidance and control, the Mengzhou Earth spacecraft has achieved autonomous rendezvous and docking with the Chinese space station, and the Mengzhou lunar spacecraft has achieved autonomous rendezvous and docking with the Lanyue spacecraft. Related Articles | Metrics

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Manned lunar landing GNC technology and verification CHEN Shangshang, LI Ji, YANG Wei, WANG Xiaolei, ZHANG Yi, ZHANG Xiaowen, WANG Pengji, GUO Chaoli, LI Yuxin, JIA Feida, XI Kun, WEN Yi 2026, 46 (2):  71-81.  doi: 10.16708/j.cnki.1000-758X.2026.0025 Abstract ( 137 )   PDF (11885KB) ( 306 )   Save The guidance, navigation, and control (GNC) technology is most critical for achieving a soft landing on the lunar surface, and its design correctness critically relies on comprehensive ground verification. Compared with unmanned missions, manned lunar landings impose higher requirements on the reliability and autonomy of the GNC system. To address these challenges, several novel methods were proposed in guidance, including predictive descent orbit strategy, range-controlled powered explicit guidance law, and optimal constant-altitude hazard avoidance. For navigation, an integrated navigation scheme was designed based on signal consistency checking technology, which combined data from inertial measurement unit, microwave radar, and optical navigation sensors. For attitude control, technologies such as rapid disturbance estimation, high-precision torque allocation, and multi-engine fault detection were developed. To verify the developed GNC technology, the Attitude and Orbit Control System (AOCS) general platform models were adopted in mathematical simulations, and model parameters for newly developed individual units were obtained through experiments. For hardware-in-the-loop simulations, a new verification system was designed, featuring capabilities such as environment simulation and manual control simulation. For full-system comprehensive verification tests, additional scenarios such as emergency ascent and main engine failure were included. The mathematical simulations achieved a landing point accuracy better than 100m, the hardware-in-the-loop simulations demonstrated a navigation error of less than 1m in 1 minute, and the emergency handling procedures in the full-system comprehensive verification tests met expectations. Guidance and control results across all three verification methods were normal. The research methodology fully inherits China's existing landing GNC technologies, develops necessary new technologies tailored to the new mission characteristics, and has been validated through simulations and testing. The research results can provide valuable references for subsequent work. Related Articles | Metrics

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Methods for improving landing stability during airbag buffering process of Mengzhou manned spacecraft WANG Guoqing, ZHU Meifang, SU Haodong, JIA He, LEI Jiangli, GONG Yiming 2026, 46 (2):  82-89.  doi: 10.16708/j.cnki.1000-758X.2026.0026 Abstract ( 94 )   PDF (5558KB) ( 183 )   Save To improve the landing stability of the Mengzhou manned spacecraft during the airbag buffering process, thereby enhancing the reliability and safety of the mission, relevant research on the underlying mechanisms and improvement methods is conducted. A finite element model (FEM) of the combined buffer airbag and the spacecraft reentry module is established based on LS-DYNA. The control volume method (CVM) is adopted to simulate the process of internal gas compression and work done during airbag buffering. Furthermore, a dynamic mechanism analysis of the airbag buffering process is conducted; combined with simulation results, the main factors affecting landing stability are further clarified. A novel overloadpressure combined control method based on differentiated active control is proposed: on the basis of the traditional overload control scheme, pressure data are incorporated, and the overturning moment generated during the buffering process is reduced through differentiated active control. Finally, the effectiveness of the proposed method is verified by the combination of multi-condition simulation comparison and full-scale model tests. The finite element simulation model of the reentry module-airbag assembly is obtained. The main factors affecting the landing stability of the reentry module during the buffering process are clarified, namely the pressure difference between the front and rear airbags and the tensile force of the airbags on the module. This is macroscopically manifested as: with other conditions fixed, the larger the noseup angle, the worse the landing stability. Three new control schemes are designed, and multi-condition comparative analysis is conducted between the traditional overload control scheme and the three new ones, followed by the selection of the optimal scheme. This scheme achieves a reduction in the maximum tilt angle of the module by 10.4%-18.2% respectively under the nose-up conditions, and the effectiveness of the new control scheme is verified through tests. An in-depth analysis is conducted on the main influencing factors and their tendencies regarding landing stability during the airbag buffering process. Based on the mechanism of the airbag buffering process, the traditional overload control scheme is optimized, which significantly improves the landing stability under various conditions. This research can provide reference for the design and optimization of landing stability in the airbag buffering system of subsequent large-scale crewed spacecraft. Related Articles | Metrics

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Design and application of communication system for landing and ascending integrated validation spacecraft LIU Yanwei, HUANG Zhen, XU Jian, SUN Xingliang, MU Chunhui 2026, 46 (2):  90-98.  doi: 10.16708/j.cnki.1000-758X.2026.0027 Abstract ( 90 )   PDF (7351KB) ( 154 )   Save In response to the stringent requirements for reliable command and telemetry transmission during ground landing and ascending experiments of the LY Lunar Lander, alongside constraints including the absence of RF telemetry and tracking link equipment and the inability to modify the ground control system's status, a wired communication system design methodology based on Time-Triggered Ethernet (TTE) and synchronous serial interfaces was proposed, aiming to high-reliability, low-latency data communication between the integrated validator and the ground during the experiment. Generalized optoelectronic conversion devices were deployed both on the spacecraft and on the ground, classifying and recognizing TTE and serial port signals, and converting them into optical signals. These signals were transmitted via fiber optic cables to establish long-distance, wired interconnection between the integrated validator platform and the ground control system. A high-speed, low-latency, multi-redundant spacecraft-to-ground communication link was established, which providing a highly reliable transmission channel for control commands, telemetry status, and various image data. After multiple ground firing tests, the spacecraft-to-ground communication system achieves a maximum rate of 100Mbit/s and data latency below 6 milliseconds. The results validate the rationality and correctness of the proposed methodology, significantly reducing development costs while ensuring the successful completion of the LY Lunar Lander's ground landing and ascending experiments. The method also offers valuable reference for the design of future lunar and Martian exploration spacecraft. Related Articles | Metrics

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Propulsion technology of the lunar module for manned lunar exploration GU Shuaihua, WANG Jianhai, HU Yuzhou, ZHU Yinjuan 2026, 46 (2):  99-107.  doi: 10.16708/j.cnki.1000-758X.2026.0028 Abstract ( 170 )   PDF (4898KB) ( 315 )   Save A series of key technical issues, such as the design of high-performance and highly reliable propulsion systems for manned lunar exploration spacecraft, lightweight high-load-bearing large-volume tanks, and throttleable engine, have been addressed. A number of concerns and systematic control measures have been proposed to solve these technical problems. The pressure and pressure difference of the storage tank are ensured by differentiating the oxygen-fuel pressure reducing valve under large system flow and by compensating for pressure drop under small flow. The thrust consistency of the four variable-thrust engines is guaranteed by balancing the upstream flow resistance of the system and the thermal calibration of individual units. The above-mentioned methods were verified through technical verification by conducting single-machine tests and system-level tests. The results show that the proposed solutions in this paper can reduce or eliminate the risks faced by various key technologies. The technical problems of pressure difference stability control of common bottom tanks under large and small flow rates and thrust consistency control of four variable-thrust engines have been solved. The pressure difference of the common bottom tank can be controlled within a small range of 0 to 0.3MPa, and the thrust deviation of the four variable-thrust engines is controlled within 100N. These researches indicate that manned lunar exploration has overcome the key technical challenges of the propulsion system, and can reliably provide power for manned lunar spacecraft, as well as offer new ideas for the design of subsequent spacecraft. Related Articles | Metrics

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Investigation of emergency return trajectory characteristics and reentry areas in crewed lunar landing missions WANG Hailiao, XU Ming, ZENG Hao, TIAN Lin, PENG Qibo 2026, 46 (2):  108-117.  doi: 10.16708/j.cnki.1000-758X.2026.0029 Abstract ( 99 )   PDF (8717KB) ( 134 )   Save One of the key distinctions between a crewed lunar mission and traditional lunar exploration missions lies in the requirement to ensure astronaut safety throughout the mission. In the event of an emergency during lunar orbitoperations, it is essential to possess a return capability that satisfies both reentry and landing constraints, which constitutes a complex Earth-Moon transfer trajectory design problem. This study employs a trajectory design method that combines a parametric approach with the solution of the Quasi-Lambert problem to systematically investigate the family characteristics and reentry regions of Earth-Moon return trajectories under such constraints. The proposed method enables the parametric generation of impulsive escape trajectories from lunar orbit at any epoch, targeting designated landing sites on Earth. On this basis, the velocity increment requirements and reachable landing regions associated with trajectory families of different transfer times are computed and analyzed. Furthermore, the feasibility and distribution characteristics of the landing sites are assessed by incorporating the declination variation patterns under different lunar phase conditions. Results demonstrate that for trajectory families with transfer times of 2, 3, and 5 days, the required velocity increments exhibit pronounced short-period and long-period coupled oscillations influenced by Earth's rotation and lunar phase evolution. Moreover, under different lunar phase declination conditions, the distribution of feasible Earth landing sites shows significant variation. To ensure emergency return capability at any epoch, the target landing site latitude should be restricted to below 20° N. If higher-latitude landing sites are desired, the lunar phase declination at the emergency epoch must fall within specific ranges. Related Articles | Metrics

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A neural network approach for fast estimation of mid-correction in trans-lunar trajectory CHANG Xiaokuan, LI Haiyang, LI Zeyue 2026, 46 (2):  118-125.  doi: 10.16708/j.cnki.1000-758X.2026.0030 Abstract ( 99 )   PDF (4556KB) ( 127 )   Save During Earth-Moon transfer, a spacecraft's orbit insertion errors propagate and amplify rapidly in a strong nonlinear dynamical environment, critically impacting mission success. Traditional mid-course correction methods, often reliant on the "small deviation" assumption and ground-based support, struggle to meet the demands for autonomous, real-time onboard execution. This study develops a lightweight neural network-based method for the rapid estimation of mid-course correction impulses, designed for limited onboard computational resources. The approach begins with high-fidelity shooting simulations and theoretical analysis, which reveal a near-linear relationship between initial velocity errors and the required correction impulses. Subsequently, a lightweight fully-connected network is constructed to establish a direct, end-to-end mapping from nominal orbital parameters and correction time to the corresponding impulse sensitivity coefficients. Validation results demonstrate that the relative error in predicting these coefficients remains below 3%. The proposed method reduces dependencies on traditional assumptions and ground support, offering a viable pathway for autonomous mid-course correction under stringent onboard resource constraints. Related Articles | Metrics

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Intelligent agent training support system for lunar scientific exploration and application DONG Zhenyuan, HE Shijing, JIANG Jiahui, LIN Qian, YANG Hanzhe, WANG Zhi, YAN Dong, LIU Chengbao, ZHANG Peng 2026, 46 (2):  126-137.  doi: 10.16708/j.cnki.1000-758X.2026.0031 Abstract ( 77 )   PDF (11631KB) ( 138 )   Save To meet the intelligent requirements of the lunar rover in complex environments, enhancing the decision-making capabilities of agents through training is of critical importance. Currently, for lunar missions there remains a lack of high-fidelity simulation environments capable of accurately reproducing complex terrain and lighting conditions. There is also an urgent need to construct diverse training data that covers extreme scenarios encountered during long-distance exploration. This achieves a joint simulation of "lunar terrain scenarios, equipment models, and environmental perception", establishing a high-fidelity training support system for lunar surface mobile intelligent agents. Firstly, utilizing digital moon technology, a centimeter-level precision lunar terrain model was constructed, providing a physically accurate foundation for agent training. And an integrated simulation scenario coupling lunar environment and rover dynamics models was established to simulate the entire workflow of lunar surface operations. Lastly, a real-time bidirectional data exchange links were established between the rover's perception, navigation, planning, and control algorithms and the simulation platform, enabling continuous learning and evolution of the intelligent agent within a closed-loop simulation. This system provides a critical technological platform for training autonomous capabilities and validating algorithms for mobile lunar rover agents, offering essential data support. It holds significant engineering value for the implementation of China's manned lunar exploration missions. Related Articles | Metrics

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A triple-reflector concentrator for lunar exploration ZHANG Yanlong, TIAN Mengfan, XIA Zhou, LI Lifang, GUO Pengzhen, DENG Zongquan 2026, 46 (2):  138-150.  doi: 10.16708/j.cnki.1000-758X.2026.0032 Abstract ( 76 )   PDF (13558KB) ( 120 )   Save To address the challenges of low solar collection efficiency, large optical aberrations, and unstable energy transmission in lunar exploration and in-situ resource utilization (ISRU), the optimization and dynamic compensation of a triple-reflector solar concentrator are investigated. The concentrator is designed with a primary mirror, a double-sided secondary mirror, and a fast steering mirror (FSM). A conjugate-correction method for secondary-mirror aberration compensation is developed, and an effective optical focality (EOF) metric is introduced to balance spherical-aberration correction and energy occlusion. Ray-tracing simulations are performed to analyze the evolution of focal-spot energy distribution and aberration patterns. In addition, a coupled dynamic model of the FSM-optical path is established based on the Lagrange formulation to realize precise angular compensation and stable beam control.The proposed system achieves a 204× concentration ratio and produces a quasi-parallel beam after complete spherical-aberration correction. EOF-based optimization improves the effective optical efficiency by approximately 15% compared with conventional dual-reflector systems. With FSM compensation, the tracking angular error is reduced to below 1 μrad, significantly enhancing the stability of long-distance optical energy transmission. The integration of the triple-reflector configuration with FSM-based control achieves coordinated optimization of high-concentration near-field focusing and collimated long-range beam output, while effectively mitigating secondary-mirror obstruction. The findings are of importance for long-term lunar missions, space-based solar power (SBSP), and high-temperature solar thermochemical applications. Related Articles | Metrics