Based on the developmental histroy and current advancements of international Mars sample return missions，an analysis of flight mission was carried out，and the necessity of Mars orbital rendezvous and the Earth-Mars transfer flight plan were discussed. On this basis，the key phases of the spacecraft for Mars sample return mission (Mars entry/descent/landing，sampling/encapsulation/transfer，Mars surface take-off and ascent，Mars orbital rendezvous，etc.) and technical difficulties were summarized. Based on typical design schemes home and abroad，the main technical solutions of the key phases were studied.

The liquid propulsion system of Mars Ascent Vehicle (MAV) provides orbital control thrust, as well as attitude control forces and torque，which is crucial for ensuring the MAV’s payload capacity, precise orbital insertion, and reduced energy requirements. The system needs to undergoe complex mission profiles, including Earth-to-Mars transfer, Mars orbital insertions, Mars EDL (Entry, Descent, and Landing) flight, and storage on the Martian surface before being operational. In this paper, the strict constraints on structural mass, envelope size, and available energy faced by the MAV’s liquid propulsion system were addressed. Key technologies required for lightweight, miniaturization, and high-precision orbit insertion, such as thrust, specific impulse, environmental adaptability and reliability, were systematically reviewed. A design solution incorporating low-freezing-point propellants, medium-chamber-pressure engines, surface tension common-bottom tanks, and a highly reliable, highly integrated system was proposed. The results provide an important reference for the formulation of Mars ascent vehicle’s design scheme.

To address the high propellant consumption during Mars orbit insertion for sample return missions，this study investigates the feasibility of utilizing Martian atmosphere through aeroassist technologies. A comprehensive review of international engineering applications and research status of aerobraking and aerocapture technologies is conducted，with in-depth analysis of key technologies including orbit strategy design，aerodynamic and aerothermodynamic environment analysis，atmospheric modeling，in-flight measurement，control strategies，as well as capture corridor design and deployable thermal protection systems. The study reveals that aerobraking is a mature technology (validated in four NASA missions)，capable of saving approximately 1.2 km/s of ΔV，with drag coefficients ranging from 1.9 to 2.0 and surface heat flux not exceeding 8.21 kW/m². Aerocapture presents challenges，particularly in deployable heatshield technology. For upcoming Mars sample return missions，aerobraking is identified as the most technically viable aeroassist approach currently，requiring emphasis on deceleration efficiency，stability，and aerothermal management with in-flight monitoring during implementation. Aerocapture necessitates dedicated efforts in optimizing capture corridors，designing specialized aerodynamic configurations，and validating deployable structures. This research provides a crucial technical reference for mission engineering.

The sample return mission is the most effective mission form for human beings to study major scientific problems in the universe. However，the scientific significance of a sample return mission is based on the effective protection of the sample from being contaminated by earth substances. Meanwhile，from the perspective of Earth safety，it is also necessary to prevent the biological risks of returning samples from celestial bodies that may have extraterrestrial life to the Earth. Therefore，strict sample protective measures need to be applied for a sample return mission. In This paper，we first outlined the sample protection requirements for a sample return mission; reviewed major sample protection measures of previous sample return missions taken by foreign countries，with a focus on the ongoing Mars 2020 mission; Next，we extracted four key technologies related to sample protection，including cleaning and sterilization technology for flight hardware，recontamination prevention technology before and after launch，sealed packaging technology for returned samples，and ground isolation and processing technology for returned samples. Further，we sorted out major sample protection works required in different phases of a sample return mission，and finally puts forward suggestions on sample protection for China’s future Mars sample return mission.

In this research, the emergency rendezvous guidance strategy in manned Mars exploration missions was analyzed and designed, and the rapid rendezvous method between Mars landing module and spacecraft when the mission is terminated during landing was mainly discussed. First, the orbit transformation process of Mars landing phase was simplified. When the mission was terminated, according to the initial phase of the landing module and the spacecraft, the height of the far point of the Mars phase-modulated orbit was determined by the special point orbit change strategy. After the phase-modulation of the landing mudole was completed, it entered the circular orbit around the Mars. Then the terminal aiming point of far-range rendezvous was designed based on CW equation. The final far-range rendezvous was realized based on the guidance strategy of the near-circular deviation equation. Finally, the proposed emergency rendezvous strategy was verified by simulation. Simulation results show that the designed emergency rendezvous guidance strategy can effectively complete the rapid rendezvous and docking between the Mars landing module and the spacecraft during the emergency rendezvous of Hohmann descent and power descent.

The launch and takeoff process of spacecraft from Mars is one of the key steps in the mission of Mars Sampling Return (MSR). Based on the Discrete Element Method (DEM)，a particle medium model of Martian soil simulant is established to study the bearing behavior of the soil. A multi-factor coupling dynamic model including Martian soil，lander，launcher and Mars Ascent Vehicle (MAV) is also constructed to perform oblique-launch simulation. Results show that the changes in attitude angles of the MAV are relatively small under different launch directions on a slope. The yaw，pitch and roll angles of the MAV are respectively no greater than 1.0°，3.2° and 0.18° during the launch and separation process. The launch process will affect the magnitude of the bearing load of the Martian soil simulant on each footpad of the lander. However，the lander’s pitch angle and sliding distance are small. The overall attitude of the lander is stable，which means the lander can effectively support the MAV’s launch process. The research method used in this paper is able to guide the dynamics simulation and stability design of planetary launch devices.

The upcoming Chinese Mars sample return missions are projected to acquire Martian subsurface samples at depths reaching 2 meters through drilling, representing the first endeavor of its kind. In this study, a comprehensive analysis of Martian deep drilling and sampling tasks was conducted, with engineering bottlenecks in subsurface excavation systematically identified. Design criteria for the drilling system were formulated according to mission requirements, leading to the proposition of a drilling and stationary sampling device prototype capable of core sampling at the depth of up to 2 meters. A dedicated drilling testbed was constructed, where verification experiments were performed using JLU Mars series soil/rock simulants. Operational parameters, drill bit geometries, and operational strategies were examined and optimized, with successful validation of the 2-meter drilling capability. The findings demonstrate the technical feasibility of the proposed solution, providing valuable insights for China’s future Mars sample return missions.

In this paper, the possible elements of space environment during Martian sample-return missions were identified and their effects on the spacecraft were analyzed. The space environment during Martian sample-return missions involves earth space and interplanetary charged particle environment, Martian atmosphere, Martian landing site surface environment and Martian surface low-pressure discharge and dust charging environment. For earth space and interplanetary charged particle environment, trapped particles in earth radiation belt, solar particle event, galactic cosmic ray and solar wind, and the effect produced by them on total ionizing dose, displacement damage dose, single event effect and surface charging effect were quantitatively analyzed. Martian atmospheric parameters were calculated using the Martian global atmosphere model, and the seasonal variations of optical depth and dust storm were analyzed based on observational data. The ground environment of potential Martian landing sites was analyzed, taking the Tianwen-1 candidate landing site as an example, with focus on the MOLA height, slope gradients, and rock distribution characteristics. In addition, special emphasis was placed on analyzing the effects of low-pressure discharge and dust-charging on the Martian surface. The results form the quantitative conditions for engineering, and can be used to support the mission design for Martian sample-return missions.

This study proposes a unified modeling and computational framework for lunar dynamics. Based on the weak-field approximation of general relativity，the framework systematically incorporates the major physical effects involved in orbital motion and dual-layer forced lunar libration，including gravitational perturbations from main-belt asteroids，the solar J₂ term，tidal deformation，and inertia tensor delay. It establishes a modular and highly consistent system by unifying dynamical derivations，reference frame definitions，and time system specifications. Using this framework，two-way Lunar Laser Ranging (LLR) data from 2015 to 2021 were processed，while the coordinates of lunar retroreflectors and terrestrial tracking stations were independently solved. The results show that the maximum deviation between the predicted orbit and the DE430 ephemeris along the Earth-Moon line is within 0.182 m，and the predicted error in physical libration Euler angles remains within 600 mas over 50 years. The post-fit root mean square (RMS) of one-way LLR residuals reaches 2.46 cm after reflector adjustment and 1.68 cm after ground station adjustment. This study demonstrates the physical consistency and practical adaptability of the unified modeling system，providing key technical support for the autonomous development of high-precision lunar ephemerides.

In this paper, a constraint satisfaction-based task planning method based on the hierarchical constraint graph was proposed for the attachment task planning of a multi-node probe. First, the representation of time-resource constraints, task-time networks, and hierarchical constraint graph models were introduced, and the planning problem was transformed into a constraint satisfaction problem. The method inferred the value range information of variables through the task-time network and employed an arc consistency algorithm with bidirectional constraint support for constraint propagation. In addition, the variable heuristic rules guided by constraint hierarchy information and value heuristic rules that prioritize resource satisfaction were designed. Experimental results demonstrate the effectiveness of the proposed method.