Mars Design Reference Architecture Research Articles

Multiphysics and multi-scale design and analysis of nuclear thermal propulsion pellet bed reactor based on spherical cermet fuel element

Due to the advantages of high thrust and high specific impulse, nuclear thermal propulsion (NTP) is a preferred near-term rocket engine technology for future Mars missions. This article introduces the reactor design based on spherical cermet fuel element, with the aim of studying the feasibility of applying this new fuel in NTP system. First, design requirements are determined based on the unique features of NTP. In terms of neutronics, the effect of rocket launch accidents and fuel loss caused by high-temperature hydrogen are considered. Thermal hydraulics refer to the requirements of NASA's Mars Design Reference Architecture 5.0 (DRA 5.0) for impulse, thrust, and thrust to weight ratio. Second, based on the requirements above, the optimal reactor configuration parameters are obtained with the goal of maximizing the thrust to weight ratio: reactor radius 51.1 cm, height 106 cm, aspect ratio 1.04, reflector 13 cm, and reactor mass 3940 kg. Finally, the performances of the optimal reactor design are analyzed. The neutron spectrum shows a characteristic of fast neutron spectrum, and fuel loss analysis shows that the core can withstand a fuel mass loss of about 4 wt%. In addition, the core axial and radial power distribution are calculated, with an overall power peak factor of 1.26. The thermal hydraulic calculation provides a range of core operating conditions that satisfy the design requirements of DRA 5.0 and temperature, with power of 620 ∼ 950 MW and flow rate of 16 ∼ 24 kg/s. Compared with the ANL200/2000 reactor using prismatic cermet, it has the advantage of higher specific impulse. The main disadvantage is the large pressure drop, which prevents it from being applied to the scenarios of high power and high thrust. Therefore, for NTP reactor using spherical cermet elements, it should be applied to low to medium power scenarios.

Shielding factors for a fission-powered mars exploration rover

Background: With the advances in compact fission power systems and Stirling converters, the efficiency and portability of electrical energy production systems has increased substantially thanks to NASA’s Kilopower project. These 1-10 kWe-class reactors are compact enough to be transportable by pressurized rovers, allowing an extended human reach across the surface of Mars to produce scientific data at a rate that greatly surpasses the rate of robotic rover data collection. If rovers were to harness fission power while mobile, it is possible for near-limitless range. This capability could allow one crew on Mars for under a “Conjunction” class mission as outlined in NASA’s Mars Design Reference Architecture (DRA) 5.0 to visit many geological sites of interest instead of requiring additional landings to explore distant features of the Martian surface. This study explored the parameters that affected the mass of the shielding required to protect the crew against a fission reactor embedded in the chassis of a pressurized rover. Methods: An analytical approach was devised to estimate the required mass of a hypothetical solid tungsten shield under various conditions. Results: Power levels below 3.4 kWe were found to be impractical for rover applications. Between 3.4 kWe and 10 kWe would be able to recharge the rover without returning to base camp. The increase in shielding mass from a 3.4 kWe reactor and a 10 kWe reactor was 6%. The reduction between a 10 kWe reactor at 3 m and 6 m was 7%. Varying the shield thickness in accordance with astronaut activities resulted in a 19% reduction. Powering the reactor off before exiting the vehicle resulted in an additional 65% reduction. Conclusions: Knowledge of the crew activities and locations had the greatest impact on the required shielding mass and the mission activities should be well-understood before finalizing a shield design.

Water Electrolysis for Propulsion of a Crewed Mars Mission

NASA’s Mars Design Reference Architecture 5.0 presents architecture options for a crewed Mars mission. This paper compares the cryogenic chemical-propulsion option in that study with an alternative: water-electrolysis propulsion. Propellant is stored as liquid water instead of cryogenic hydrogen and oxygen, and then electrolyzed on demand into gaseous hydrogen and oxygen for combustion. This addresses a technology gap in the reference architecture: cryogenic propellant storage life. Water is inert and stable, allowing indefinite storage. The mission refuels from pre-positioned tanks at a lunar libration point and Mars orbit, significantly reducing vehicle mass. Low-thrust transfers between Earth escape, Mars encounter, and vice versa reduce the need to store gas for impulsive maneuvers. Only relatively small impulsive maneuvers are used for orbit injection, escape, and plane changes. The proposed architecture achieves the same mission with at most the same number of launch vehicles and without cryogenic propellant storage. Both architectures use five superheavy lift launch vehicles; however, the reference architecture demands all five within 120 days. The alternative is more flexible. Pre-positioning propellant ahead of time allows a less demanding launch cadence. Potential utilization of water from the moon or elsewhere reduces the number of launches from five to as few as two.

Generalized Multicommodity Network Flow Model for the Earth–Moon–Mars Logistics System

Simple logistics strategies such as “carry-along” and Earth-based “resupply” were sufficient for past human space programs. Next-generation space logistics paradigms are expected to be more complex, involving multiple exploration destinations and in situ resource utilization. Optional in situ resource utilization brings additional complexity to the interplanetary supply chain network design problem. This paper presents an interdependent network flow modeling method for determining optimal logistics strategies for space exploration and its application to the human exploration of Mars. It is found that a strategy using lunar resources in the cislunar network may improve overall launch mass to low Earth orbit for recurring missions to Mars compared to NASA’s Mars Design Reference Architecture 5.0, even when including the mass of the in situ resource utilization infrastructures that need to be predeployed. Other findings suggest that chemical propulsion using liquid oxygen/liquid hydrogen, lunar in situ resource utilization water production, and the use of aerocapture significantly contribute to reducing launch mass from Earth. A sensitivity analysis of in situ resource utilization reveals that, under the given assumptions, local lunar resources become attractive at productivity levels above in the context of future human exploration of Mars.

Technology development for human exploration of Mars

Current plans call for the first human missions to Mars to be launched perhaps as early as 2035. The recently completed “Mars Design Reference Architecture 5.0” study defines a conceptual mission architecture and identifies enabling technologies. NASA is beginning long range development on key technologies needed for these missions because it will take many years for them to reach maturity. The ISS and the lunar outpost will be used as test beds for these technologies to reduce risk and prepare for human exploration of Mars. NASA’s Exploration Technology Development Program is maturing technologies and demonstrating operational scenarios for lunar exploration that are extensible to future human missions to Mars. These include entry, descent, and landing systems for large payloads; fission surface power systems; liquid oxygen–liquid methane propulsion systems; cryogenic fluid management; closed-loop life support; small pressurized rovers for surface mobility; in-situ resource utilization; radiation shielding; and optical communications. Advanced technologies will enable more affordable and sustainable Mars exploration.