Design Reference Architecture Research Articles

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 and taken into account for shield design.

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.

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

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.

Stay Cool—Alternatives for Long-Term Storage of Large Quantities of Liquid Hydrogen on a Mars Transfer Vehicle

Improved methods for storing liquid hydrogen in larger quantities and over longer periods of time in space are becoming progressively more critical as sights are once again set on Mars. Current storage methods involve the venting of vaporized hydrogen to space, with the consequence that significant amounts of hydrogen are wasted. Extra hydrogen must be stored to account for this loss resulting in unnecessary mass penalties. Eliminating this waste can reduce overall mission mass, extend mission range, and perhaps most importantly, lower mission trip times and costs. This technical note explores alternative methods for storing liquid hydrogen with an emphasis on missions to Mars as laid out by the National Aeronautics and Space Administration’s current design reference architecture.

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.

Seeded Hydrogen in Mars Transfer Vehicles Using Nuclear Thermal Propulsion Engines

Using the NASA Design Reference Architecture, an Aerojet Rocketdyne vehicle architecture, and a University of Alabama in Huntsville nuclear thermal propulsion engine model, the performance of a Mars transfer vehicle using hydrogen propellant seeded with argon was analyzed. Seeded hydrogen, up to a maximum seed mass concentration of 55.85%, increased engine and vehicle performance by reducing pressure losses, decreasing reactor power, and increasing the overall while assuming constant propellant volume and vehicle dry mass. The tradeoff for using seeded hydrogen was lowered specific impulse and increased net propellant mass, resulting in increased vehicle wetted mass. Vehicle performance monotonically increased with seed mass concentration and provided a best-case 32-day reduction in round-trip transit time to Mars versus pure hydrogen.

Numerical Investigation and Parametric Study on Thermal-Hydraulic Characteristics of Particle Bed Reactors for Nuclear Thermal Propulsion

Nuclear thermal propulsion (NTP) could be an advanced technology to facilitate a new and excellent rocket engine that would at least double the performance of the best conventional chemical rocket engines. NTP has been under development for several decades and was selected as the leading candidate technique for the manned mission to Mars, as suggested in Design Reference Architecture 5.0. During development, many concepts have been proposed, designed, and tested. Among which, the particle bed reactor (PBR) is the one of highest performance, and its compact and lightweight features make it ideal for space applications. In this paper, the thermal-hydraulic characteristics of a PBR are mainly investigated through two studies. The first study is to evaluate whether the principles derived from the PBR of uniform heat release could be applied in the cases of a nonuniform heating profile. The second study is to analyze the effects of some aspects, including porosity of the hot frit and cold frit, power shift, inlet temperature of the coolant on the internal flow, and heat transfer processes in the PBR of a nonuniform heat release. These findings may provide technical support for the subsequent design and optimization of the PBR.

Seeded Hydrogen in Nuclear Thermal Propulsion Engines

Using a NASA Design Reference Architecture, a nuclear thermal propulsion engine model was developed to numerically examine the effects on engine performance by adding heavy noble gases into the hydrogen propellant stream (seeding) before the reactor inlet. Seeded hydrogen, up to a maximum seed mass concentration of 55.85%, increased engine performance by reducing pressure losses, decreasing reactor power, and improving convective heat transfer. Seeding hydrogen also has the potential to increase the engine thrust by up to 13.75% by using the decreased pressure losses. The decreased reactor power reduces the cool down propellant mass of the vehicle, whereas the increased bulk density of the propellant can provide more overall. The tradeoff for introducing a seed into the hydrogen stream is lowered specific impulse.

Mass Optimization of a Supercritical CO2 Brayton Cycle Power Conversion System for a Mars Surface Fission Power Reactor

In the National Aeronautics and Space Administration (NASA) Design Reference Architecture 5.0 (DRA 5.0), fission surface power systems (FSPS) are described as “enabling for the human exploration of Mars.” This study investigates the design of a power conversion system (PCS) based on supercritical carbon dioxide (sCO2) Brayton configurations for a growing Martian colony. Various configurations utilizing regeneration, intercooling (IC), and reheating are analyzed. A model to estimate the mass of the PCS is developed and used to obtain a realistic mass-optimized configuration. This mass model is conservative, being based on simple concentric tube counterflow heat exchangers and published data regarding turbomachinery masses. For load following and redundancy purposes, the FSPS consists of three 333 kWe reactors and PCS to provide a total of 1 MWe for 15 years. The optimal configuration is a sCO2 Brayton cycle with 60% regeneration and two stages of intercooling. The majority of the analyses are performed in matlab, with certain data provided by a comsol multiphysics model of part of a low-enriched uranium (LEU) ceramic metallic (CERMET) reactor core.

Passive Reactivity Control of Nuclear Thermal Propulsion Reactors

This paper explores the possibility of passively controlling the reactivity of a nuclear thermal propulsion (NTP) reactor. The objective of this study is to limit the use of the radial control drums to start-up and shutdown procedures and ensure that the exact same operation is performed for each full-power burn. To achieve the goal, this work considers several design measures, which include a low-density burnable absorber in the tie-tube components of the core, the use of variable hydrogen density in the moderator element coolant passages, and the judicious selection of a modified mission profile to maximize the decay of 135Xe after operation. In addition, the improved stability from the enhanced fuel temperature feedback due to the implementation of low-enriched-uranium fuel is also exploited for the realization of passive reactivity control. In this work, a passive reactivity control system is implemented in the Superb Use of Low Enriched Uranium (SULEU) NTP core and analyzed in terms of its ability to fulfill a NASA Mars Mission Design Reference Architecture 5.0–style mission. It is concluded that the use of the control drums can be limited to start-up and shutdown operations only, eliminating operator input in order to maintain a constant power level in the core.

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.

Prototype Development of an Integrated Mars Atmosphere and Soil-Processing System

The concept of living off the land by using the indigenous resources of the Moon, Mars, or other potential sites of robotic and human exploration has been termed in situ resource utilization (ISRU) and will be an enabling technology to open up the solar system. Although the most recent National Aeronautics and Space Administration (NASA) human Mars mission study (Design Reference Architecture 5.0) showed that production of propellants and life support consumables was a mission-enabling capability, mission planners were hesitant to select the newly proposed water extraction from Mars soil option because of the perceived high risk associated with this approach. To overcome resistance in putting ISRU capabilities in the critical path of mission success, NASA ISRU developers have adopted the approach of designing and building hardware into end-to-end systems at representative mission scales and testing these systems under mission-relevant conditions at analog field test sites. Previous ISRU field demonstrations have been standalone lunar ISRU modules running on alternating current power with nonoptimal integration. The primary goal of the Mars atmosphere and regolith collector/processor for lander operations (MARCO POLO) project is to design, build, and test an end-to-end first-generation Mars ISRU atmospheric and soil-processing system powered by mission-relevant direct current power while also demonstrating closed-loop power production via the combination of a fuel cell and electrolyzer. A secondary goal is to perform remote and autonomous operations with this integrated system on a 3×3-m(∼9×9-ft) octagon lander and transfer oxygen and methane produced to a cryocart for use with a thruster to demonstrate an end-to-end Mars resource-to-thrust concept. This paper will outline the overall design, technologies used, and concept of operations for the MARCO POLO project and its upcoming field demonstration.

Reference Architecture for Personalized E-learning Systems using Proxy and Caching (RAPESPAC)

Service-Oriented approach is the most appropriate architectural approach that is widely used for building highly large-scale enterprise systems. SORAPES is a serviceoriented reference architecture designed for enterprise elearning systems. This architecture is designed by re-using web services and learning objects for distributed environment. It is notified that the system generate more traffic during authentication/ authorization process and retrieval of information from the database. The problem here is to minimize the traffic during authentication and retrieval ode learning objects process. This traffic will be minimized by centrally managing caching scheme and secured web service requests. This improves the performance and scalability of SORAPES in distributed environment, and hence it is proposed to design reference architecture using proxy and caching to authentication. The objective of this research paper is to design a reference architecture using proxy and caching called RAPESPAC. The RAPESPAC reduces the load on large-scale e-learning applications in distributed environment by centrally managing and caching the web service requests. The important features of RAPESPAC are load balancing, proxy, caching and policy. This architecture could be deployed with confidence in high transaction volume service environments providing high availability of data and reliable security mechanism. General Terms Service-Oriented Computing, E-learning Systems, Distributed Computing.

Simplified scenario for manned Mars missions

We propose a simplified but efficient scenario for a manned Mars mission. The idea is to select a crew of only 2 astronauts and to bring in situ resource utilization systems in the same vehicle. For security reasons, we suggest duplicating the mission as it was proposed by Von Braun. At very moment of the journey, the two vehicles would stay close so that each crew could provide help to the other. We show that this scenario is much simpler than the last design reference architecture proposed by NASA. The initial mass in low Earth orbit is minimized and the risks are also reduced. The total cost could be in the order of 40 billion dollars.

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.