Comparison of Four Space Propulsion Methods for Reducing Transfer Times of Crewed Mars Mission

We assess the possibility of reducing the travel time of a crewed mission to Mars by examining four different propulsion methods and keeping the mass at departure under 2500 tonne, for a fixed architecture. We evaluated representative systems of three different state of the art technologies (chemical, nuclear thermal and electric) and one advance technology, the ``Pure Electro-Magnetic Thrust’’ (PEMT) concept (proposed by Rubbia). A mission architecture mostly based on the Design Reference Architecture 5.0 is assumed in order to estimate the mass budget, that influences the performance of the propulsion system. Pareto curves of the duration of the mission and time of flight versus mass of mission are drawn. We conclude that the ion engine technology, combined with the classical chemical engine, yields the shortest mission times for this architecture with the lowest mass and that chemical propulsion alone is the best to minimise travel time. The results obtained using the PEMT suggest that it could be a more suitable solution for farther destinations than Mars.

💡 Research Summary

The paper investigates how to shorten the total travel time of a crewed mission to Mars by evaluating four distinct propulsion technologies while keeping the launch mass below 2 500 tonnes. The authors adopt a mission architecture based on the Design Reference Architecture 5.0 (DRA 5.0), which includes a crew of six, a cargo spacecraft, a crew transfer vehicle, and a descent/ascent vehicle. The baseline DRA 5.0 mission uses a minimum‑energy Hohmann‑type transfer that takes about 180 days each way, plus roughly 500 days of surface stay, leading to a total mission duration of about 900 days.

Four propulsion concepts are selected as representative of current state‑of‑the‑art or near‑future technologies:

1. Common Extensible Cryogenic Engine (CECE) – a high‑thrust, moderate‑specific‑impulse (≈450 s) chemical engine using liquid hydrogen/oxygen. It can be restarted many times, making it suitable for impulsive burns such as departure, Mars capture, and landing.

2. Nuclear Engine for Rocket Vehicle Application II (NERVA II) – a nuclear thermal rocket with Isp≈800 s and thrust≈34 kN, providing higher efficiency than chemical propulsion but requiring a heavy reactor, shielding, and cooling system.

3. Radio‑Frequency Ion Technology XT (RIT‑XT) – an electric ion thruster powered by solar arrays. It delivers very high Isp (≈4 600 s) but low thrust (≈0.12 N/kg). The mass of the solar panels and power‑control electronics is explicitly accounted for.

4. Pure Electro‑Magnetic Thrust (PEMT) – a speculative concept proposed by Carlo Rubbia that converts mass directly into energy via nuclear reactions and uses the resulting photons for thrust. In theory it offers extremely high thrust‑to‑mass efficiency, but practical implementation demands massive radiation shielding and heat‑rejection hardware.

For each technology the authors construct scaling relationships between engine size, thrust, specific impulse, and system mass, then generate Pareto curves that relate mission duration (both total and transfer‑only) to launch mass. They consider both impulsive (chemical, nuclear) and continuous (electric, PEMT) thrust profiles, and they model the required propellant, reactor fuel, solar‑array area, and ancillary hardware for each case.

Key Findings

• Chemical‑only (CECE) provides the absolute shortest transfer time (≈180 days) because its high thrust minimizes gravity losses, but the required propellant mass pushes the total launch mass close to the 2 500‑ton limit, leaving little margin for payload or redundancy.

• Hybrid Chemical + Electric (CECE for departure/arrival burns, RIT‑XT for the cruise phase) achieves a transfer time of roughly 150 days while keeping the total launch mass around 2 300 tonnes. The electric stage dramatically reduces propellant consumption, and the chemical stage supplies the high thrust needed for orbital insertion and landing.

• Nuclear Thermal (NERVA II) yields a modest improvement over pure chemical (≈160 days) but the reactor, shielding, and coolant add significant dry mass, so the overall launch mass is again near the upper bound.

• PEMT can theoretically reduce the cruise to about 130 days, but the mass of the required radiation shield, heat‑radiator, and the large power‑conversion hardware inflates the system mass to >3 000 tonnes, violating the study’s mass constraint. Consequently, PEMT is not advantageous for a Mars mission under current assumptions, though its mass‑to‑energy conversion could become attractive for missions to more distant destinations where solar power is weak.

The authors also discuss secondary considerations such as the impact of propulsion choice on mission risk, system complexity, and the need for technologies like in‑situ resource utilization (ISRU) or aerobraking. They argue that while high‑thrust impulsive engines minimize travel time, the associated propellant penalty makes them less practical for large‑scale crewed missions. Conversely, low‑thrust electric propulsion reduces propellant but requires large power‑generation systems, increasing dry mass and potentially affecting reliability.

Conclusions

• The most realistic and efficient solution for a crewed Mars mission, given current technology and a 2 500‑ton launch mass ceiling, is a hybrid architecture that pairs a high‑thrust chemical engine for impulsive maneuvers with a high‑specific‑impulse electric ion thruster for the interplanetary cruise.

• Pure chemical propulsion remains the fastest in terms of raw travel time but is mass‑inefficient.

• Nuclear thermal propulsion offers a middle ground but does not provide a clear advantage over the hybrid approach when system mass is accounted for.

• PEMT, while conceptually promising, is presently unsuitable for Mars missions due to its heavy ancillary hardware; however, it may become a viable option for outer‑planet or interstellar precursor missions once the shielding and thermal‑management challenges are solved.

The paper recommends further research on improving PEMT shielding, advancing nuclear‑thermal reactor designs with lighter materials, and optimizing solar‑array efficiency for electric propulsion, to broaden the feasible propulsion portfolio for future deep‑space crewed missions.