The Mars Sample Return (MSR) mission is a major but prohibitively expensive goal of planetary exploration. This article presents a conceptual single-launch integrated mission architecture that consolidates entry, descent, and landing (EDL), surface sample retrieval, Mars ascent, and Earth-return functions (traditionally distributed across three separate spacecraft) into one capsule-class vehicle delivered by one heavy-lift expendable launcher. Here, “single-launch” applies only to a single launch from Earth to the Moon. The subsequent ascent from the Martian surface in the Mars Ascent–Earth Return Vehicle (MA-ERV) is an internal phase of the mission and not a separate launch campaign. The architecture consists of a three-stage Mars Ascent–Earth Return Vehicle (MA-ERV) composed of liquid-bipropellant and solid-propellant chemical stages coupled with a Hall-effect electric propulsion stage for low-thrust spiral escape. Patch-conic trajectory analysis for the 2029 primary and 2031 backup launch windows and pork chop plot evaluation and interplanetary return trajectory trade study (ballistic coast vs. continued electric propulsion thrusting) with gravity loss and drag corrections results in a total mission ΔV of 9. 14 km/s. According to Edelbaum’s low-thrust approximation, with validation against numerical propagation of orbits (2%–5% discrepancy), within the wider literature of electric propulsion trajectory optimization, eclipse, J 2 perturbation, and thrust degradation sensitivity analysis, prediction of Mars escape occurs in approximately 270 days (85% thrust duty cycle) using 49 kg Xe at 5. 0 kW input power. A three-degrees-of-freedom Monte Carlo EDL simulation (10, 000 cases), complemented by qualitative six-degrees-of-freedom sensitivity analysis and parametric thermal protection system (TPS) mass sizing, results in a 99th-percentile landing ellipse of 8. 7 km × 4. 2 km with peak heat flux within PICA-class thermal protection limits. A physics-based four-legged robot energy model with penalties for Martian regolith interaction, cross-validated against terrestrial data from the Spot platform, predicts 6. 2-h long sorties with a 4. 6 km operational radius using in-situ resource utilization (ISRU) produced methane–oxygen fuel cells in 40% system efficiency. Mass estimation using AIAA S-120A-compliant, subsystem-level mass growth allowances and lognormal probabilistic analysis results in a total launch mass of approximately 18, 200 kg using current best estimates, which is some 8% above the Falcon Heavy expendable capacity of 16, 800 kg. This outcome highlights mass budget closure as the primary technical issue and encourages the development of higher capability launch vehicles (such as Falcon Heavy with performance improvements, Vulcan Centaur Heavy, or New Glenn) or subsystem mass reduction. Sensitivity and risk analyses confirm the feasibility of the conceptual architecture under nominal conditions but show that subsystem-level mass growth characteristic of technology readiness level (TRL) 3–4 technologies makes closure on the baseline Falcon Heavy launcher with no mitigation impossible. A preliminary planetary protection compliance assessment shows the Earth Return Vehicle containment architecture as a critical design driver that requires 420 kg allocation. The proposed architecture has an estimated mission cost of 3. 0–6. 0 billion, which is a potential cost reduction compared with similar remaining elements of the current campaign by the National Aeronautics and Space Administration/European Space Age (NASA/ESA) Mars Surveyor (MSR) campaign (estimated to be 5–8 billion without accounting for the already-operational Perseverance rover). This study identifies a mass-critical but physically plausible pathway for near-term Mars sample return at the conceptual design level while acknowledging that significant technology maturation and high fidelity design validation and advance to Phase A requires a possible launcher upgrade.
Wang Xing (Thu,) studied this question.