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NASA NTRS2007

Design of a Resistively Heated Thermal Hydraulic Simulator for Nuclear Rocket Reactor Cores

A preliminary design study is presented for a non-nuclear test facility which uses ohmic heating to replicate the thermal hydraulic characteristics of solid core nuclear reactor fuel element passages. The basis for this testing capability is a recently commissioned nuclear thermal rocket environments simulator, which uses a high-power, multi-gas, wall-stabilized constricted arc-heater to produce high-temperature pressurized hydrogen flows representative of reactor core environments, excepting radiation effects. Initially, the baseline test fixture for this non-nuclear environments simulator was configured for long duration hot hydrogen exposure of small cylindrical material specimens as a low cost means of evaluating material compatibility. It became evident, however, that additional functionality enhancements were needed to permit a critical examination of thermal hydraulic effects in fuel element passages. Thus, a design configuration was conceived whereby a short tubular material specimen, representing a fuel element passage segment, is surrounded by a backside resistive tungsten heater element and mounted within a self-contained module that inserts directly into the baseline test fixture assembly. With this configuration, it becomes possible to create an inward directed radial thermal gradient within the tubular material specimen such that the wall-to-gas heat flux characteristics of a typical fuel element passage are effectively simulated. The results of a preliminary engineering study for this innovative concept are fully summarized, including high-fidelity multi-physics thermal hydraulic simulations and detailed design features.

Litchford, Ron J.↗

NASA NTRSDate not supplied

Recent Advancements in Electrical Capacitance Mass Gauging for Cryogenic PropellantTanks

The current lack of unsettled mass gauging is a key roadblock for many space activities, such as orbital refueling, missions to the Moon and Mars, and nuclear thermal and nuclear electric propulsion technologies. Liquid can form any one of an infinite number of configurations in microgravity, such as floating in globs or accumulating on tank surfaces in discontiguous volumes, or both. Capacitive sensing requires no moving parts and dissipates close to zero heat, making it an ideal candidate for cryogenic fluid mass gauging in settled and unsettled configurations. Capacitive sensing has a history of proven use in space propellant tanks, including tanks on the space shuttle, Saturn V, and the Apollo Lunar Excursion Module. Modern capacitive sensor technology allows using the entire tank as the capacitor volume by placing electrodes on the tank walls and propellant management surfaces. Capacitance is directly related to density, and therefore to mass for fixed volumes. In this presentation, we discuss the concepts behind whole-tank capacitance mass gauging and associated engineering challenges. We describe recent efforts to develop a micro-g unsettled cryogenic mass gauge using whole-tank capacitance sensing, including the development of test beds, electronics, and algorithms. We describe several mathematical processing techniques, including empirical-based averaging, electrical capacitance volume tomography, and spatial regularization. A modeling study, performed using settled configurations in gravity and no gravity, and with a set of 100 random fluid configurations, has indicated that spatial regularization, in which capacitance measurements are weighted to account for non-uniform electric fields, yields a mass fraction accuracy of 8% for any fluid configuration. Such a sensor is expected to operate in real time with a sampling frequency of at least 1 kHz.

cryogenic↗

NASA NTRS1972

Completely modular thermionic reactor ion propulsion system /trips/.

Description of a nuclear reactor powered ion propulsion system which is an advanced completely modularized system lending itself to development of prototype and/or flight type components without the need for complete system tests until late in the development program. This modularity is achieved in all of the subsystems and components of the electric propulsion system, including the thermionic fuel elements, the heat rejection subsystem (head pipes), the power conditioning modules, and the ion thrusters. The use of heat pipes for cooling the collectors and as a means of heat transport to the radiator allows early prototype or flight configuration testing of a small module of the heat rejection subsystem as opposed to full scale liquid metal pumps and radiators in a large vacuum chamber. The power conditioner (p/c) is arranged in modules with passive cooling which allows complete prototype testing. The ion engines are typically matched with one or more p/c modules and are the same size for any power level propulsion system of interest.

Peelgren, M. L.↗

NASA NTRS2015

Reliability and Probabilistic Risk Assessment - How They Play Together

Since the Space Shuttle Challenger accident in 1986, NASA has extensively used probabilistic analysis methods to assess, understand, and communicate the risk of space launch vehicles. Probabilistic Risk Assessment (PRA), used in the nuclear industry, is one of the probabilistic analysis methods NASA utilizes to assess Loss of Mission (LOM) and Loss of Crew (LOC) risk for launch vehicles. PRA is a system scenario based risk assessment that uses a combination of fault trees, event trees, event sequence diagrams, and probability distributions to analyze the risk of a system, a process, or an activity. It is a process designed to answer three basic questions: 1) what can go wrong that would lead to loss or degraded performance (i.e., scenarios involving undesired consequences of interest), 2) how likely is it (probabilities), and 3) what is the severity of the degradation (consequences). Since the Challenger accident, PRA has been used in supporting decisions regarding safety upgrades for launch vehicles. Another area that was given a lot of emphasis at NASA after the Challenger accident is reliability engineering. Reliability engineering has been a critical design function at NASA since the early Apollo days. However, after the Challenger accident, quantitative reliability analysis and reliability predictions were given more scrutiny because of their importance in understanding failure mechanism and quantifying the probability of failure, which are key elements in resolving technical issues, performing design trades, and implementing design improvements. Although PRA and reliability are both probabilistic in nature and, in some cases, use the same tools, they are two different activities. Specifically, reliability engineering is a broad design discipline that deals with loss of function and helps understand failure mechanism and improve component and system design. PRA is a system scenario based risk assessment process intended to assess the risk scenarios that could lead to a major/top undesirable system event, and to identify those scenarios that are high-risk drivers. PRA output is critical to support risk informed decisions concerning system design. This paper describes the PRA process and the reliability engineering discipline in detail. It discusses their differences and similarities and how they work together as complementary analyses to support the design and risk assessment processes. Lessons learned, applications, and case studies in both areas are also discussed in the paper to demonstrate and explain these differences and similarities.

Safie, Fayssal↗

NASA NTRS2014

Affordable Development and Optimization of CERMET Fuels for NTP Ground Testing

CERMET fuel materials for Nuclear Thermal Propulsion (NTP) are currently being developed at NASA's Marshall Space Flight Center. The work is part of NASA's Advanced Space Exploration Systems Nuclear Cryogenic Propulsion Stage (NCPS) Project. The goal of the FY12-14 project is to address critical NTP technology challenges and programmatic issues to establish confidence in the affordability and viability of an NTP system. A key enabling technology for an NCPS system is the fabrication of a stable high temperature nuclear fuel form. Although much of the technology was demonstrated during previous programs, there are currently no qualified fuel materials or processes. The work at MSFC is focused on developing critical materials and process technologies for manufacturing robust, full-scale CERMET fuels. Prototypical samples are being fabricated and tested in flowing hot hydrogen to understand processing and performance relationships. As part of this initial demonstration task, a final full scale element test will be performed to validate robust designs. The next phase of the project will focus on continued development and optimization of the fuel materials to enable future ground testing. The purpose of this paper is to provide a detailed overview of the CERMET fuel materials development plan. The overall CERMET fuel development path is shown in Figure 2. The activities begin prior to ATP for a ground reactor or engine system test and include materials and process optimization, hot hydrogen screening, material property testing, and irradiation testing. The goal of the development is to increase the maturity of the fuel form and reduce risk. One of the main accomplishmens of the current AES FY12-14 project was to develop dedicated laboratories at MSFC for the fabrication and testing of full length fuel elements. This capability will enable affordable, near term development and optimization of the CERMET fuels for future ground testing. Figure 2 provides a timeline of the development and optimization tasks for the AES FY15-17 follow on program.

Hickman, Robert R.↗

NASA NTRS2017

Chandra Reveals Heavy Obscuration and Circumnuclear Star Formation in Seyfert 2 Galaxy NGC 4968

We present the Chandra imaging and spectral analysis of NGC 4968, a nearby (z = 0.00986) Seyfert 2 galaxy. We discover extended (approx. 1 kpc) X-ray emission in the soft band (0.5-2 keV) that is neither coincident with the narrow line region nor the extended radio emission. Based on spectral modeling, it is linked to on-going star formation [approx. 2.6-4 Mass compared to Earth yr(exp.− 1)]. The soft emission at circumnuclear scales (inner approx. 400 pc) originates from hot gas, with kT approx. 0.7 keV, while the most extended thermal emission is cooler (kT approx. 0.3 keV). We refine previous measurements of the extreme Fe K alpha equivalent width in this source (EW 2.5 + 2.6/-1.0 keV), which suggests the central engine is completely embedded within Compton-thick levels of obscuration. Using physically motivated models fit to the Chandra spectrum, we derive a Compton-thick column density [N(sub H) is greater than 1.25× 10(exp 24) cm(exp.− 2)] and an intrinsic hard (2-10 keV) X-ray luminosity of approx. 3-8× 10(exp. 42) erg s(exp. − 1) (depending on the presumed geometry of the obscurer), which is over two orders of magnitude larger than that observed. The large Fe K Alpha EW suggests a spherical covering geometry, which could be confirmed with X-ray measurements above 10 keV. NGC 4968 is similar to other active galaxies that exhibit extreme Fe K Alpha EWs (i.e., greater than 2 keV) in that they also contain on-going star formation. This work supports the idea that gas associated with nuclear star formation may increase the covering factor of the enshrouding gas and play a role in obscuring active galactic nuclei.

LaMassa, Stephanie M.↗

NASA NTRS2014

Advanced Compatibility Characterization Of AF-M315E With Spacecraft Propulsion System Materials Project

All spacecraft require propulsion systems for thrust and maneuvering. Propulsion systems can be chemical, nuclear, electrical, cold gas or combinations thereof. Chemical propulsion has proven to be the most reliable technology since the deployment of launch vehicles. Performance, storability, and handling are three important aspects of liquid chemical propulsion. Bipropellant systems require a fuel and an oxidizer for propulsion, but monopropellants only require a fuel and a catalyst for propulsion and are therefore simpler and lighter. Hydrazine is the state of the art propellant for monopropellant systems, but has drawbacks because it is highly hazardous to human health, which requires extensive care in handling, complex ground ops due to safety and environmental considerations, and lengthy turnaround times for reusable spacecraft. All users of hydrazine monopropellant must contend with these issues and their associated costs. The development of a new monopropellant, intended to replace hydrazine, has been in progress for years. This project will apply advanced techniques to characterize the engineering properties of materials used in AF-M315E propulsion systems after propellant exposure. AF-M315E monopropellant has been selected HQ's Green Propellant Infusion Mission (GPIM) to replace toxic hydrazine for improved performance and reduce safety and health issues that will shorten reusable spacecraft turn-around time. In addition, this project will fundamentally strengthen JSC's core competency to evaluate, use and infuse liquid propellant systems.

McClure, Mark B.↗

NASA NTRS2018

Electrically Heated Testing of the Kilowatt Reactor Using Stirling TechnologY (KRUSTY) Experiment Using a Depleted Uranium Core

The Kilopower project aims to develop and demonstrate scalable fission-based power technology for systems capable of delivering 1 to 10 kW of electric power with a specific power ranging from 2.5 to 6.5 W/kg. This technology could enable high-power science missions or could be used to provide surface power for manned missions to the Moon or Mars. NASA has partnered with the U.S. Department of Energy's National Nuclear Security Administration, Los Alamos National Laboratory, Nevada National Security Site (NNSS), and Y−12 National Security Complex to develop and test a prototypic reactor and power system using existing facilities and infrastructure. This technology demonstration, referred to as the "Kilowatt Reactor Using Stirling TechnologY (KRUSTY)," will undergo nuclear ground testing by the end of calendar year (CY) 2017 at the NNSS. The 1-kWe variation of the Kilopower system was chosen for the KRUSTY demonstration. The concept for the 1-kWe flight system consists of a 4 kWt highly enriched uranium-molybdenum reactor operating at 800 degC coupled to sodium heat pipes. The heat pipes deliver heat to the hot ends of eight 125-W Stirling convertors producing a net electrical output of 1 kW. Waste heat is rejected using titanium-water heat pipes coupled to carbon composite radiator panels. The KRUSTY test, based on this design, uses a prototypic highly enriched uranium-molybdenum core coupled to prototypic sodium heat pipes. The heat pipes transfer heat to two Advanced Stirling Convertors (ASC−E2s) and six thermal simulators, which simulate the thermal draw of full-scale power conversion units. Thermal simulators and Stirling engines are gas cooled. The most recent project milestone was the completion of nonnuclear system-level testing using an electrically heated depleted uranium (DU) (nonfissioning) reactor core simulator at the NASA Glenn Research Center. System-level testing has validated performance predictions and has demonstrated system-level operation and control in a test configuration that replicates the one to be used at the Device Assembly Facility (DAF) at the NNSS. Fabrication, assembly, and testing of the DU core has allowed for higher fidelity system-level testing at Glenn, and has validated the fabrication methods to be used on the highly enriched uranium core that will supply heat for the DAF KRUSTY demonstration.

Briggs, Maxwell H.↗

NASA NTRS2019

Thermal, Fluid and Neutronic Analysis of an LEU Nuclear Thermal Propulsion Core

This paper describes the use of detailed multidisciplinary fluid/thermal/ structural/neutronic simulations to predict performance of the nuclear fuel elements of a Nuclear Thermal Propulsion rocket reactor. To achieve maximum performance, a rocket reactor's fuel must operate near thermal hydraulic, structural and neutronic limits where multidisciplinary interactions are important. Yet physical testing is expensive, time- consuming and risky. Lower-fidelity correlations (heat transfer) and simulations have always existed for design, and one role of detailed numerical analysis is to confirm correlation validity and accuracy. For complex and subtle issues, detailed numerical simulations may prove their value. The paper gives examples of both of these situations. Limitations of the methods and potential extensions will be explored.

Rocket Engine Design↗

NASA NTRS1996

High Leverage Space Transportation System Technologies for Human Exploration Missions to the Moon and Beyond

The feasibility of returning humans to the Moon by 2004, the 35th anniversary of the Apollo 11 landing, is examined assuming the use of existing launch vehicles (the Space Shuttle and Titan 4B), a near term, advanced technology space transportation system, and extraterrestrial propellant--specifically 'lunar-derived' liquid oxygen or LUNOX. The lunar transportation system (LTS) elements consist of an expendable, nuclear thermal rocket (NTR)-powered translunar injection (TLI) stage and a combination lunar lander/Earth return vehicle (LERV) using cryogenic liquid oxygen and hydrogen (LOX/LH2) chemical propulsion. The 'wet' LERV, carrying a crew of 2, is configured to fit within the Shuttle orbiter cargo bay and requires only modest assembly in low Earth orbit. After Earth orbit rendezvous and docking of the LERV with the Titan 4B-launched NTR TLI stage, the initial mass in low Earth orbit (IMLEO) is approx. 40 t. To maximize mission performance at minimum mass, the LERV carries no return LOX but uses approx. 7 t of LUNOX to 'reoxidize' itself for a 'direct return' flight to Earth followed by an 'Apollo-style' capsule recovery. Without LUNOX, mission capability is constrained and the total LTS mass approaches the combined Shuttle-Titan 4B IMLEO limit of approx. 45 t even with enhanced NTR and chemical engine performance. Key technologies are discussed, lunar mission scenarios described, and LTS vehicle designs and characteristics are presented. Mission versatility provided by using a small 'all LH2' NTR engine or a 'LOX-augmented' derivative, either individually or in clusters, for outer planet robotic orbiter, small Mars cargo, lunar 'commuter', and human Mars exploration class missions is also briefly discussed.

Borowski, Stanley K.↗

NASA NTRS1991

Optimum rocket propulsion for energy-limited transfer

In order to effect large-scale return of extraterrestrial resources to Earth orbit, it is desirable to optimize the propulsion system to maximize the mass of payload returned per unit energy expended. This optimization problem is different from the conventional rocket propulsion optimization. A rocket propulsion system consists of an energy source plus reaction mass. In a conventional chemical rocket, the energy source and the reaction mass are the same. For the transportation system required, however, the best system performance is achieved if the reaction mass used is from a locally available source. In general, the energy source and the reaction mass will be separate. One such rocket system is the nuclear thermal rocket, in which the energy source is a reactor and the reaction mass a fluid which is heated by the reactor and exhausted. Another energy-limited rocket system is the hydrogen/oxygen rocket where H2/O2 fuel is produced by electrolysis of water using a solar array or a nuclear reactor. The problem is to choose the optimum specific impulse (or equivalently exhaust velocity) to minimize the amount of energy required to produce a given mission delta-v in the payload. The somewhat surprising result is that the optimum specific impulse is not the maximum possible value, but is proportional to the mission delta-v. In general terms, at the beginning of the mission it is optimum to use a very low specific impulse and expend a lot of reaction mass, since this is the most energy efficient way to transfer momentum. However, as the mission progresses, it becomes important to minimize the amount of reaction mass expelled, since energy is wasted moving the reaction mass. Thus, the optimum specific impulse will increase with the mission delta-v. Optimum I(sub sp) is derived for maximum payload return per energy expended for both the case of fixed and variable I(sub sp) engines. Sample missions analyzed include return of water payloads from the moons of Mars and of Saturn.

Zuppero, Anthony↗

NASA NTRS2017

Simulation-Based Height of Burst Map for Asteroid Airburst Damage Prediction

Entry and breakup models predict that airburst in the Earth's atmosphere is likely for asteroids up to approximately 200 meters in diameter. Objects of this size can deposit over 250 megatons of energy into the atmosphere. Fast-running ground damage prediction codes for such events rely heavily upon methods developed from nuclear weapons research to estimate the damage potential for an airburst at altitude. (Collins, 2005; Mathias, 2017; Hills and Goda, 1993). In particular, these tools rely upon the powerful yield scaling laws developed for point-source blasts that are used in conjunction with a Height of Burst (HOB) map to predict ground damage for an airburst of a specific energy at a given altitude. While this approach works extremely well for yields as large as tens of megatons, it becomes less accurate as yields increase to the hundreds of megatons potentially released by larger airburst events. This study revisits the assumptions underlying this approach and shows how atmospheric buoyancy becomes important as yield increases beyond a few megatons. We then use large-scale three-dimensional simulations to construct numerically generated height of burst maps that are appropriate at the higher energy levels associated with the entry of asteroids with diameters of hundreds of meters. These numerically generated HOB maps can then be incorporated into engineering methods for damage prediction, significantly improving their accuracy for asteroids with diameters greater than 80-100 m.

Burst Map↗

NASA NTRS2003

Development Efforts Expanded in Ion Propulsion: Ion Thrusters Developed With Higher Power Levels

The NASA Glenn Research Center was the major contributor of 2-kW-class ion thruster technology to the Deep Space 1 mission, which was successfully completed in early 2002. Recently, NASA s Office of Space Science awarded approximately $21 million to Glenn to develop higher power xenon ion propulsion systems for large flagship missions such as outer planet explorers and sample return missions. The project, referred to as NASA's Evolutionary Xenon Thruster (NEXT), is a logical follow-on to the ion propulsion system demonstrated on Deep Space 1. The propulsion system power level for NEXT is expected to be as high as 25 kW, incorporating multiple ion thrusters, each capable of being throttled over a 1- to 6-kW power range. To date, engineering model thrusters have been developed, and performance and plume diagnostics are now being documented. The project team-Glenn, the Jet Propulsion Laboratory, General Dynamics, Boeing Electron Dynamic Devices, the Applied Physics Laboratory, the University of Michigan, and Colorado State University-is in the process of developing hardware for a ground demonstration of the NEXT propulsion system, which comprises a xenon feed system, controllers, multiple thrusters, and power processors. The development program also will include life assessments by tests and analyses, single-string tests of ion thrusters and power systems, and finally, multistring thruster system tests in calendar year 2005. In addition, NASA's Office of Space Science selected Glenn to lead the development of a 25-kW xenon thruster to enable NASA to conduct future missions to the outer planets of Jupiter and beyond, under the High Power Electric Propulsion (HiPEP) program. The development of a 100-kW-class ion propulsion system and power conversion systems are critical components to enable future nuclear-electric propulsion systems. In fiscal year 2003, a team composed of Glenn, the Boeing Company, General Dynamics, the Applied Physics Laboratory, the Naval Research Laboratory, the University of Wisconsin, the University of Michigan, and Colorado State University will perform a 6-month study that will result in the design of a 25-kW ion thruster, a propellant feed system, and a power processing architecture. The following 2 years will involve hardware development, wear tests, single-string tests of the thruster-power circuits and the xenon feed system, and subsystem service life analyses. The 2-kW-class ion propulsion technology developed for the Deep Space 1 mission will be used for NASA's discovery mission Dawn, which involves maneuvering a spacecraft to survey the asteroids Ceres and Vesta. The 6-kW-class ion thruster subsystem technology under NEXT is scheduled to be flight ready by calendar year 2006. The less mature 25- kW ion thruster system under HiPEP is expected to be ready for a flight advanced development program in calendar year 2006.

Patterson, Michael J.↗

NASA NTRS2015

NASA's Nuclear Thermal Propulsion Project

Space fission power systems can provide a power rich environment anywhere in the solar system, independent of available sunlight. Space fission propulsion offers the potential for enabling rapid, affordable access to any point in the solar system. One type of space fission propulsion is Nuclear Thermal Propulsion (NTP). NTP systems operate by using a fission reactor to heat hydrogen to very high temperature (>2500 K) and expanding the hot hydrogen through a supersonic nozzle. First generation NTP systems are designed to have an Isp of approximately 900 s. The high Isp of NTP enables rapid crew transfer to destinations such as Mars, and can also help reduce mission cost, improve logistics (fewer launches), and provide other benefits. However, for NTP systems to be utilized they must be affordable and viable to develop. NASA's Advanced Exploration Systems (AES) NTP project is a technology development project that will help assess the affordability and viability of NTP. Early work has included fabrication of representative graphite composite fuel element segments, coating of representative graphite composite fuel element segments, fabrication of representative cermet fuel element segments, and testing of fuel element segments in the Compact Fuel Element Environmental Tester (CFEET). Near-term activities will include testing approximately 16" fuel element segments in the Nuclear Thermal Rocket Element Environmental Simulator (NTREES), and ongoing research into improving fuel microstructure and coatings. In addition to recapturing fuels technology, affordable development, qualification, and utilization strategies must be devised. Options such as using low-enriched uranium (LEU) instead of highly-enriched uranium (HEU) are being assessed, although that option requires development of a key technology before it can be applied to NTP in the thrust range of interest. Ground test facilities will be required, especially if NTP is to be used in conjunction with high value or crewed missions. There are potential options for either modifying existing facilities or constructing new ground test facilities. At least three potential options exist for reducing (or eliminating) the release of radioactivity into the environment during ground testing. These include fully containing the NTP exhaust during the ground test, scrubbing the exhaust, or utilizing an existing borehole at the Nevada National Security Site (NNSS) to filter the exhaust. Finally, the project is considering the potential for an early flight demonstration of an engine very similar to one that could be used to support human Mars or other ambitious missions. The flight demonstration could be an important step towards the eventual utilization of NTP.

Houts, Michael↗

NASA NTRS1993

Power Conditioning System Modelling for Nuclear Electric Propulsion

NASA LeRC is currently developing a Fortran based model of a complete nuclear electric propulsion (NEP) vehicle that would be used for piloted and cargo missions to the Moon or Mars. The proposed vehicle design will use either a Brayton or K-Rankine power conversion cycle to drive a turbine coupled with a rotary alternator. Two thruster types are also being studied, ion and magnetoplasmadynamic (MPD). In support of this NEP model, Rocketdyne developed a power management and distribution (PMAD) subroutine that provides parametric outputs for selected alternator operating voltages and frequencies, thruster types, system power levels, and electronics coldplate temperatures. The end-to-end PMAD model described is based on the direct use of the alternator voltage and frequency for transmitting power to either ion or MPD thrusters. This low frequency transmission approach was compared with dc and high frequency ac designs, and determined to have the lowest mass, highest efficiency, highest reliability and lowest development costs. While its power quality is not as good as that provided by a high frequency system, it was considered adequate for both ion and MPD engine applications. The low frequency architecture will be used as the reference in future NEP PMAD studies.

Metcalf, Kenneth J.↗

NASA NTRS2024

NASA Marshall Space Flight Center In-Space Cryogenic Propulsion Capabilities and Applications to Human Exploration

The current focus on lunar exploration and future human missions to Mars is driving in-space propulsion system requirements toward higher performance cryogenic systems with long-duration storage and operational capabilities. Not only do these systems offer higher performance than storable propellant options, but they also enable the potential for in-situ propellant production. Future Mars transit systems are envisioned to utilize either high-thrust nuclear thermal propulsion (with liquid hydrogen propellant), or hybrid systems with both cryogenic chemical systems (likely LOX/CH4) for high acceleration maneuvers and nuclear electric systems for long duration high Isp maneuvers. Exploration architectures based on either of these options require the use of high-performance cryogenic propellants with long-duration storage capabilities for both in-space transportation as well as planetary descent and ascent functions. Current efforts focusing on lunar exploration also rely on cryogenic propellants (either LOX/LCH4 or LOX/LH2) for lunar transit and descent/ascent transportation functions. In-space cryogenic propulsion systems pose numerous technology challenges with respect to long-duration propellant storage and usage, including advanced insulation, tank stratification and pressure management, cryogenic refrigeration to reduce propellant loss through boil off, low leakage cryogenic valves, low temperature liquid acquisition, and cryogenic propellant transfer. NASA has invested in technology development efforts, demonstrating individual technologies and systems-level operations. NASA Marshal Space Flight Center has also invested in multiple test facilities and modular test rigs that allow ground demonstration of numerous integrated technologies and systems concepts of operations. Additional investments have been made to mature analytical capabilities and design tools. These capabilities, both test/demonstration & engineering design/analysis, are available to support internal efforts and industrial partners in the development of exploration and science mission systems. With this increased interest, it is critical to understand the current state of in-space cryogenic propulsion technology, determine risks to its successful application to human exploration, and prepare the engineering, test, and evaluation capabilities to support the ambitious plans for future systems. This paper provides a survey of recent developments in in-space cryogenic propulsion and cryogenic propellant management technologies, as well as facilities and engineering/analytical capabilities ready to support current and future exploration efforts.

Cryogenic propulsion, cryogenic fluid management, ↗

NASA NTRS2024

NASA Marshall Space Flight Center In-Space Cryogenic Propulsion Capabilities and Applications to Human Exploration

Cryogenic propulsion, cryogenic fluid management, ↗

NASA NTRS2024

NASA Marshall Space Flight Center In-Space Cryogenic Propulsion Capabilities and Applications to Human Exploration

The current focus on lunar exploration and future human missions to Mars is driving in-space propulsion system requirements toward higher performance cryogenic systems with long-duration storage and operational capabilities. Not only do these systems offer higher performance than storable propellant options, but they also enable the potential for in-situ propellant production. Future Mars transit systems are envisioned to utilize either high-thrust nuclear thermal propulsion (with liquid hydrogen propellant), or hybrid systems with both cryogenic chemical systems (likely LOX/CH 4 ) for high acceleration manoeuvres and nuclear electric systems for long duration high Isp manoeuvres. Exploration architectures based on either of these options require the use of high-performance cryogenic propellants with long-duration storage capabilities for both in-space transportation as well as planetary descent and ascent functions. Current efforts focusing on lunar exploration also rely on cryogenic propellants (either LOX/LCH 4 or LOX/LH 2 ) for lunar transit and descent/ascent transportation functions. In-space cryogenic propulsion systems pose numerous technology challenges with respect to long-duration propellant storage and usage, including advanced insulation, tank stratification and pressure management, cryogenic refrigeration to reduce propellant loss through boil off, low leakage cryogenic valves, low temperature liquid acquisition, and cryogenic propellant transfer. NASA has invested in technology development efforts, demonstrating individual technologies and systems-level operations. NASA Marshal Space Flight Center has also invested in multiple test facilities and modular test rigs that allow ground demonstration of numerous integrated technologies and systems concepts of operations. Additional investments have been made to mature analytical capabilities and design tools. These capabilities (both test/demonstration & engineering design/analysis) are available to support both internal efforts and industrial partners in the development of exploration and science mission systems. With this increased interest, it is critical to understand the current state of in-space cryogenic propulsion technology, determine risks to its successful application to human exploration, and prepare the engineering, test, and evaluation capabilities to support the ambitious plans for future systems. This paper provides a survey of recent developments in in-space cryogenic propulsion and cryogenic propellant management technologies, as well as facilities and engineering/analytical capabilities ready to support current and future exploration efforts.

Cryogenic propulsion↗