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| Fig. 1: NASA concept illustration of a nuclear electric propulsion spacecraft above Mars. (Courtesy of NASA. Source: Wikimedia Commons) |
For future missions to the Moon, Mars, and eventually other planets, it is necessary to develop alternative spacecraft propulsion solutions beyond traditional chemical rockets. The Mars Transportation Assessment Study (MTAS) has proposed the development of a nuclear-propulsion spacecraft for missions to Mars starting in the early 2040s. [1] However, an independent review by the National Academies of Sciences, Engineering, and Medicine (NASEM) indicates that current technological maturity is insufficient to meet these objectives. [2] The report emphasizes the need for substantial and immediate investment, particularly to overcome the operational lifetime limitations of propulsion systems. [3]
A crewed mission to Mars goes beyond the limits of engineering. The first major challenge is the total velocity change required, as an Earth-to-Mars transfer demands a ΔV of approximately 6 to 10 km/s, depending on mission architecture. The second challenge involves meeting lifetime requirements for missions lasting several months, given that even a minimum-energy transfer takes about 259 days one way. Finally, there is the large mass that must be launched and accelerated out of Earth's gravity. Thus, Mars transportation is months of deep-space operations.
Today, chemical propulsion remains the leading technology, as it offers high thrust and is a mature solution. However, for Mars missions, the propellant logistics, staging, and operational complexity grow rapidly, especially when large payload deliveries are required.
Nuclear propulsion is compelling because it reshapes these trades in two distinct ways. Nuclear Thermal Propulsion (NTP) can preserve rocket-like thrust while roughly doubling propellant efficiency. [4] Nuclear Electric Propulsion (NEP) can reach far higher exhaust velocities than chemical systems, but it becomes fundamentally constrained by available electrical power, heat rejection, and thruster lifetime. [3]
The reference human mission is a two-year mission: the round-trip time from Earth orbit departure to Earth orbit capture is 760 days, with a 30 sol surface stay for the first mission. The propulsion requirement is expressed as total ΔV, and it varies strongly with the Earth-Mars opportunity, which repeats about every 26 months. For consistency in the initial theoretical comparison, we adopt a fixed ΔV of 9 km/s as a conservative reference case. This ΔV starts from a high Earth orbit, after launch and assembly, and therefore does not include Earth launch or Mars entry.
With ΔV fixed, we now focus on specific impulse (Isp), which measures how efficiently a propulsion system uses propellant. It is defined as
where F is thrust, ṁ is propellant mass flow rate, and g0 is standard gravity. For a given ΔV requirement, a higher Isp reduces the propellant mass needed, which directly reduces the vehicle's total mass.
To connect the required ΔV to the vehicle's non-propellant mass, we use the Tsiolkovsky rocket equation,
where m0 is the initial mass before the burn and mf is the final mass after propellant is expended. Rearranging gives mf/m0 = exp(-ΔV/(g0Isp)). For the reference case ΔV = 9 km/s, this yields the following ideal mass ratios and corresponding payload fractions:
Chemical (Isp = 450 s): 13.0%.
NTP (Isp = 900 s): 36.1%.
NEP (Isp = 2600 s): 70.3%.
These payload fractions are upper bounds, because in a real vehicle mf is not payload only: it also includes tanks, structure, engines/thrusters, power conversion, radiators, and radiation shielding.
For a more realistic chemical stage estimate, a simple first-order assumption is that, for the propulsion system mass excluding payload, about 10% is hardware (tanks, structure, and engine), defined relative to the combined mass of hardware plus propellant at the start of the burn. With Isp = 450 s and Δ V = 9 km/s, this yields a practical payload fraction of about 3.4%, which is lower than the 13% ideal upper bound that assumes the final mass is payload only. We therefore use 3.4% as the reference payload fraction for chemical propulsion in the comparisons that follow.
Nuclear propulsion can be divided into two main families.
Nuclear Electric Propulsion (NEP), as illustrated in Fig. 1, uses a reactor as an electrical power plant, and electric thrusters accelerate propellant to high exhaust velocities. The added value is extraordinary propellant efficiency, with Isp typically on the order of 2,000 to 10,000 s, but the thrust is limited by available electrical power and by the efficiency and lifetime of the thruster system.
NEP is attractive for Mars because it targets the main problem of interplanetary missions: carrying propellant. High exhaust velocity means a spacecraft can accumulate large total impulse using far less propellant mass than chemical propulsion. In a typical NEP architecture, the main subsystems include a fission reactor, radiation shielding, a power conversion system (turning reactor heat into electricity), power conditioning and distribution (to regulate and route power to the thrusters), propellant storage, and clusters of electric thrusters. [1]
The Space Nuclear Propulsion project at NASA's Marshall Space Flight Center is currently evaluating two potential thrusters for NEP: the Hall effect thruster (HET), which uses an electric field to accelerate ions while a magnetic field traps electrons, and the magnetoplasmadynamic (MPD) thruster, which accelerates an ionized gas using the electromagnetic force generated by high current in a magnetic field. [3] The MPD thruster presents several performance advantages that could be decisive compared to the Hall effect thruster. For instance, the power per thruster, for an optimal MPD operation, can far exceed the performance of HETs. [3] In a Mars mission scenario, using an MPD thruster may reduce the spacecraft's minimum launch mass from 519 tons (HETs) to 486 tons. [2]
The key difficulty is that neither HETs nor MPD thrusters have been demonstrated at Mars-relevant power/current levels for Mars-relevant durations. A Mars NEP transfer can imply many thousands of hours of cumulative thrusting, whereas high-power thrusters have limited long-duration test heritage at the relevant scales. A few-kW Hall thruster can surpass 10,000 hours on the ground, but the same lifetime at hundreds of kilowatts has not been demonstrated yet.
A concrete mass breakdown helps quantify the payload fraction once shielding and power hardware are included. For the COMPASS 2039 mission, a 760-day round-trip Mars mission using a NEP spacecraft with Isp = 2600 s, the initial mass in low Earth orbit is 469.3 t, with a 45.3 t payload. This corresponds to a payload fraction of 45.3/469.3 ≈ 9.7%. [5] This illustrates the key tradeoff: high Isp reduces required propellant, but the power system and shielding create a substantial non-propellant mass penalty that directly lowers the useful payload fraction.
NTP addresses Mars transportation from the opposite direction. Unlike NEP, the goal here is not to trade thrust for efficiency. Instead, NTP aims to preserve high thrust, enabling short burns and flexible trajectories while increasing exhaust velocity relative to chemical engines.
NTP uses a reactor as a high-temperature heat source. As in a rocket engine, the propellant, often hydrogen, is heated and then expanded through a nozzle. The added value relative to chemical propulsion is a higher exhaust velocity: about 8.8 km/s for Isp ≈ 900 s, compared to 3.4-4.4 km/s for chemical rockets (Isp ≈ 350-450 s). [4] However, an Isp of 900 s corresponds to a hydrogen reactor exit temperature on the order of 2700 K, illustrating why materials and reactor fuel elements are central engineering drivers. [2]
The choice of hydrogen as a propellant increases the complexity of an NTP spacecraft because it is cryogenic and difficult to store for long periods without boil-off management. Thus, NTP can reduce propellant mass for interplanetary burns while increasing complexity in cryogenic storage and thermal control.
A representative NTP mass breakdown is available from NASAs Mars Design Reference Architecture 5.0 Copernicus vehicle concept. For the 2033 mission (Isp = 900s), the reported initial mass in low Earth orbit is 488.5 t with a 59.8 t payload. This corresponds to a payload fraction of 59.8/488.5 ≈ 12.2%. The NTP stage mass includes additional external radiation shielding for crew protection. [6]
Nuclear propulsion reshapes the Mars mission design trade-offs and introduces a new optimization problem. Compared with chemical architectures, where feasibility is often driven by propellant fraction and the operational burden of moving and storing large amounts of propellant, nuclear concepts tend to be driven by shielding geometry, reactor and power-conversion mass, radiator size, and the ability to demonstrate long-duration, reliable operation.
Within nuclear options, NEP and NTP address Mars transportation in fundamentally different ways. NEP leverages extremely high exhaust velocity to reduce propellant mass and can be especially attractive for missions that demand large total impulse, such as cargo delivery and high-ΔV transfers. NTP preserves rocket thrust and enables flexible trajectories, with improved propellant efficiency relative to chemical propulsion. Its principal challenges are concentrated in high temperature reactor operation and cryogenic hydrogen storage.
As a system-level benchmark, published architecture studies show that the useful payload fraction can vary widely once power systems, shielding, and trajectory assumptions are included. In the NEP case, the reported mass breakdown corresponds to a useful payload fraction of ~9.7%, while in the NTP case the implied useful payload fraction is ~12.2%. For reference, our chemical stage sized to the fixed ΔV = 9 km/s theoretical comparison yields ~3.4% payload fraction. Because these case studies use different launch/assembly orbits, mission opportunities, transfer types, and ΔV, these percentages should be read as illustrative system-level points rather than directly comparable performance metrics. Still, they highlight the core trade: nuclear propulsion can increase payload fraction relative to chemical architectures, but the added system complexity can reduce the mass advantage.
Both NEP and NTP can enable compelling Mars missions, but each requires targeted development and system-level demonstration before it can be considered credible for human Mars transportation. [2]
© Allan Attia. The author warrants that the work is the author's own and that Stanford University provided no input other than typesetting and referencing guidelines. The author grants permission to copy, distribute and display this work in unaltered form, with attribution to the author, for noncommercial purposes only. All other rights, including commercial rights, are reserved to the author.
[1] "Mars Transportation Assessment Study", U.S. National Aeronautics and Space Administration, March 2023.
[2] Space Nuclear Propulsion for Human Mars Exploration (National Academies Press, 2021).
[3] L. S. Mason et al., "Nuclear Power Concepts and Development Strategies for High-Power Electric Propulsion Missions to Mars", U.S. National Aeronautics and Space Administration, NASA/TM-20210016968, March 2022.
[4] R. H. Frisbee, "Advanced Space Propulsion for the 21st Century", Journal of Propulsion and Power 19, 1129 (2003).
[5] S. R. Oleson et al., "Compass Final Report: Nuclear Electric Propulsion (NEP) - Chemical Vehicle 1.2", NASA Glenn Research Center, NASA/TM-20210017131, September 2021
[6] S. K. Borowski, D. R. McCurdy, and T. W. Packard, "Modular Growth NTR Space Transportation System for Future NASA Human Lunar, NEA and Mars Exploration Missions", American Insitute of Aeronautics and Astronautics, AIAA-2012-5144, 8 Sep 12.
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