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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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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 NTRS2010

Integration of In-Situ Resource Utilization Into Lunar/Mars Exploration Through Field Analogs

The NASA project to develop In-Situ Resource Utilization (ISRU) technologies, in partnership with commercial and international collaborators, has achieved full system demonstrations of oxygen production using native regolith simulants. These demonstrations included robotic extraction of material from the terrain, sealed encapsulation of material in a pressurized reactor; chemical extraction of oxygen from the material in the form of water, and the electrolysis of water into oxygen and hydrogen for storage and reuse. These successes have provided growing confidence in the prospects of ISRU oxygen production as a credible source for critical mission consumables in preparation for and during crewed missions to the moon and other destinations. Other ISRU processes, especially relevant to early lunar exploration scenarios, have also been shown to be practical, including the extraction of subsurface volatiles, especially water, and the thermal processing of surface materials for civil engineering uses and for thermal energy storage. This paper describes these recent achievements and current NASA ISRU development and demonstration activity. The ability to extract and process resources at the site of exploration into useful products such as propellants, life support and power system consumables; and radiation and rocket exhaust plume debris shielding, known as In-Situ Resource Utilization or ISRU, has the potential to significantly reduce the launch mass, risk, and cost of robotic and human exploration of space. The incorporation of ISRU into missions can also significantly influence technology selection and system development in other areas such as power, life support, and propulsion. For example. the ability to extract or produce large amounts of oxygen and/or water in-situ could minimize the need to completely close life support air and water processing system cycles, change thermal and radiation protection of habitats, and influence propellant selection for ascent vehicles and surface propulsive hoppers. While concepts and even laboratory work on evaluating and developing ISRU techniques such as oxygen extraction from lunar regolith have been going on since before the Apollo 11 Moon landing, no ISRU system has ever flown in space, and only recently have ISRU technologies been developed at a scale and at a system level that is relevant to actual robotic and human mission applications. Because ISRU hardware and systems have never been demonstrated or utilized before on robotic or human missions, architecture and mission planners and surface system hardware developers are hesitant to rely on ISRU products and services that are critical to mission and system implementation success. To build confidence in ISRU systems for future missions and assess how ISRU systems can best influence and integrate with other surface system elements, NASA, with international partners, are performing analog field tests to understand how to take advantage of ISRU capabilities and benefits with the minimum of risk associated with introducing this game-changing approach to exploration. This paper will describe and review the results of four analog field tests (Moses Lake in 6/08, Mauna Kea in 11/08. Flagstaff in 9/09; and Mauna Kea in 1/10) that have begun the process of integrating ISRU into robotic and human exploration systems and missions, and propose future ISRU-related analog field test activities that can be performed in collaboration with international space agencies.

Sanders, Gerald B.↗
NASA NTRSDate not supplied

Integrated In-Situ Resource Utilization Modeling of a Lunar Water Processing System

A key element of achieving a sustained surface presence, such as defined in NASA’s Artemis plan, is In-Situ Resource Utilization (ISRU). ISRU is the practice of using local resources to provide mission consumables that reduce system launch mass requirements, and regenerate resources (chiefly, water and oxygen) for propulsion and life support supporting both Lunar and Martian missions. ISRU systems require multiple complex processes, such as excavation, chemical reactors, and electrolysis subsystems that must operate in harmony to optimize the overall system process from beginning to end. The Mission Analysis and Integration Tool (MAIT) connects individual subsystem models into a customized, flexible framework for the purpose of technology downselect, optimization, and end-to-end process planning. MATLAB was favorable to use as the main software integration tool due to its ability to communicate with a vast number of other programming languages and makes up the backbone of data flow between inputs and outputs to the subsystem models. MAIT initially evaluated a suite of technologies, including but not limited to, the water processing Lunar Auger Dryer for ISRU (LADI) system and an oxygen extraction carbothermal reduction process. With individual models consolidated, the MAIT tool generated over 60,000 cases during its parametric sweeps; these system iterations produced valuable insight into the optimal LADI geometry for minimizing heater energy demands, estimating carbothermal reactor and radiator mass relationships, and calculated the power dynamics of the electrolysis unit. Efforts to update the MAIT tool are ongoing to accommodate and scale ISRU technologies supporting the Space Technology Mission Directorate’s (STMD) commercialization strategy, and increase the MAIT software capability to handle a wide array of ISRU system models beyond the Lunar environment, e.g. production of propellant for a Martian lander.

Avery Carlson↗
NASA NTRS2017

Robust Exploration and Commercial Missions to the Moon Using LANTR Propulsion and Lunar Liquid Oxygen Derived from FeO-Rich Pyroclastic Deposits

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. It is a demonstrated technology capable of generating both high thrust and high specific impulse (Isp approx.900 s) twice that of todays best chemical rockets. Nuclear lunar transfer vehicles consisting of a propulsion stage using three approx.16.5 klbf Small Nuclear Rocket Engines (SNREs), an in-line propellant tank, plus the payload can enable a variety of reusable lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong tourism missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing a robust in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The use of lunar liquid oxygen (LLO2) derived from iron oxide (FeO)-rich volcanic glass beads, found in numerous pyroclastic deposits on the Moon, can significantly reduce the launch mass requirements from Earth by enabling reusable, surface-based lunar landing vehicles (LLVs) using liquid oxygen/hydrogen (LO2/H2) chemical rocket engines. Afterwards, a LO2/H2 propellant depot can be established in lunar equatorial orbit to supply the LTS. At this point a modified version of the conventional NTR called the LOX-augmented NTR, or LANTR is introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an afterburner into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engines choked sonic throat - essentially scramjet propulsion in reverse. By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and Isp values while the reactor core power level remains relatively constant. A LANTR-based LTS offers unique mission capabilities including short transit time crewed cargo transports. Even a commuter shuttle service may be possible allowing one-way trip times to and from the Moon on the order of 36 hours or less. If only 1 of the extracted LLO2 propellant from identified resource sites were available for use in lunar orbit, such a supply could support daily commuter flights to the Moon for many thousands of years! The proposed paper outlines an evolutionary architecture and examines a variety of mission types and transfer vehicle designs, along with the increasing demands on LLO2 production as mission complexity and (Delta)V requirements increase. A comparison of vehicle features and engine operating characteristics, for both NTR and LANTR engines, is also provided along with a discussion of the propellant production and mining requirements associated with using FeO-rich volcanic glass as source material.

Nuclear rocket engines↗
NASA NTRS2011

U.S. Space Radioisotope Power Systems and Applications: Past, Present and Future

Radioisotope power systems (RPS) have been essential to the U.S. exploration of outer space. RPS have two primary uses: electrical power and thermal power. To provide electrical power, the RPS uses the heat produced by the natural decay of a radioisotope (e.g., plutonium-238 in U.S. RPS) to drive a converter (e.g., thermoelectric elements or Stirling linear alternator). As a thermal power source the heat is conducted to whatever component on the spacecraft needs to be kept warm; this heat can be produced by a radioisotope heater unit (RHU) or by using the excess heat of a radioisotope thermoelectric generator (RTG). As of 2010, the U.S. has launched 41 RTGs on 26 space systems. These space systems have ranged from navigational satellites to challenging outer planet missions such as Pioneer 10/11, Voyager 1/2, Galileo, Ulysses, Cassini and the New Horizons mission to Pluto. In the fall of 2011, NASA plans to launch the Mars Science Laboratory (MSL) that will employ the new Multi-Mission Radioisotope Thermoelectric Generator (MMRTG) as the principal power source. Hundreds of radioisotope heater units (RHUs) have been launched to provide warmth to Apollo 11, used to provide heating of critical components in a seismic experiment package, Pioneer 10/11, Voyager 1/2, Galileo, Cassini, Mars Pathfinder, MER rovers, etc. to provide temperature control to critical spacecraft electronics and other mechanical devices such as propulsion system propellant valves. A radioisotope (electrical) power source or system (RPS) consists of three basic elements: (1) the radioisotope heat source that provides the thermal power, (2) the converter that transforms the thermal power into electrical power and (3) the heat rejection radiator. Figure 1 illustrates the basic features of an RPS. The idea of a radioisotope power source follows closely after the early investigations of radioactivity by researchers such as Henri Becquerel (1852-1908), Marie Curie (1867-1935), Pierre Curie (1859-1906) and R. J. Strut. Almost 100 years ago, in 1913, English physicist H. G. J. Moseley (1887-1915) constructed the first nuclear battery using a vacuum flask and 20 mCi of radium (Corliss and Harvey, 1964, Proceedings of the Royal Society, 1913). After World War II, serious interest in radioisotope power systems in the U.S. was sparked by studies of space satellites such as North American Aviation s 1947 report on nuclear space power and the RAND Corporation s 1949 report on radioisotope power. (Greenfield, 1947, Gendler and Kock, 1949). Radioisotopes were also considered in early studies of nuclear-powered aircraft (Corliss and Harvey, 1964). In 1951, the U.S. Atomic Energy Commission (AEC) signed several contracts to study a 1-kWe space power plant using reactors or radioisotopes. Several of these studies, which were completed in 1952, recommended the use of RPS. (Corliss and Harvey, 1964). In 1954, the RAND Corporation issued the summary report of the Project Feedback military satellite study in which radioisotope power was considered (Lipp and Salter, 1954, RAND). Paralleling these studies, in 1954, K. C. Jordan and J. H. Birden of the AEC s Mound Laboratory conceived and built the first RTG using chromel-constantan thermocouples and a polonium-210 (210Po or Po-210) radioisotope heat source (see Figure 2). While the power produced (1.8 mWe) was low by today s standards, this first RTG showed the feasibility of RPS. A second thermal battery was built with more Po-210, producing 9.4 mWe. Jordan and Birden concluded that the Po-210 thermal battery would have about ten times the energy of ordinary dry cells of the same mass (Jordan and Birden, 1954). The heat source consisted of a 1-cm-diameter sphere of 57 Ci (1.8 Wt) of Po-210 inside a capsule of nickel-coated cold-rolled steel all inside a container of Lucite. The thermocouples were silver-soldered chromel-constantan. The thermal battery produced 1.8 mWe.

Cataldo, Robert L.↗
NASA NTRS2017

Robust Exploration and Commercial Missions to the Moon Using LANTR Propulsion and In-Situ Propellants Derived from Lunar Polar Ice (LPI) Deposits

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. It is a demonstrated technology capable of generating both high thrust and high specific impulse (Isp 900 s) twice that of todays best chemical rockets. Nuclear lunar transfer vehicles consisting of a propulsion stage using three approx.16.5 klbf "Small Nuclear Rocket Engines (SNREs)", an in-line propellant tank, plus the payload can enable a variety of reusable lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong "tourism" missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing a robust in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The processing of LPI deposits (estimated to be approx. 2 billion metric tons) for propellant production - specifically liquid oxygen (LO2) and hydrogen (LH2) can significantly reduce the launch mass requirements from Earth and can enable reusable, surface-based lunar landing vehicles (LLVs) using LO2/LH2 chemical rocket engines. Afterwards, LO2/LH2 propellant depots can be established in lunar polar and equatorial orbits to supply the LTS. At this point a modified version of the conventional NTR called the LO2-augmented NTR, or LANTR would be introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants (LDPs) for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an afterburner into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engines choked sonic throat essentially scramjet propulsion in reverse. By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and Isp values while the reactor core power level remains relatively constant. A LANTR-based LTS offers unique mission capabilities including short transit time crewed cargo transports. Even a commuter shuttle service may be possible allowing one-way trip times to and from the Moon on the order of 36 hours or less. If only 1 of the postulated water ice trapped in deep shadowed craters at the lunar poles were available for use in lunar orbit, such a supply could support daily commuter flights to the Moon for many thousands of years! The proposed paper outlines an evolutionary mission architecture and examines a variety of mission types and transfer vehicle designs, along with the increasing demands on LDP production as mission complexity and delta V requirements increase. A comparison of vehicle features and engine operating characteristics are also provided together with a discussion of the propellant production and mining requirements, and issues, associated with using LPI as the source material.

Nuclear Thermal Propulsion↗
NASA NTRS2004

Main Power Distribution Unit for the Jupiter Icy Moons Orbiter (JIMO)

Around the year 2011, the Jupiter Icy Moons Orbiter (JIMO) will be launched and on its way to orbit three of Jupiter s planet-sized moons. The mission goals for the JIMO project revolve heavily around gathering scientific data concerning ingredients we, as humans, consider essential: water, energy and necessary chemical elements. The JIM0 is an ambitious mission which will implore propulsion from an ION thruster powered by a nuclear fission reactor. Glenn Research Center is responsible for the development of the dynamic power conversion, power management and distribution, heat rejection and ION thrusters. The first test phase for the JIM0 program concerns the High Power AC Power Management and Distribution (PMAD) Test Bed. The goal of this testing is to support electrical performance verification of the power systems. The test bed will incorporate a 2kW Brayton Rotating Unit (BRU) to simulate the nuclear reactor as well as two ION thrusters. The first module of the PMAD Test Bed to be designed is the Main Power Distribution Unit (MPDU) which relays the power input to the various propulsion systems and scientific instruments. The MPDU involves circuitry design as well as mechanical design to determine the placement of the components. The MPDU consists of fourteen relays of four different variations used to convert the input power into the appropriate power output. The three phase system uses 400 Vo1ts(sub L-L) rms at 1000 Hertz. The power is relayed through the circuit and distributed to the scientific instruments, the ION thrusters and other controlled systems. The mechanical design requires the components to be positioned for easy electrical wiring as well as allowing adequate room for the main buss bars, individual circuit boards connected to each component and power supplies. To accomplish creating a suitable design, AutoCAD was used as a drafting tool. By showing a visual layout of the components, it is easy to see where there is extra room or where the components may interfere with one another. By working with the electrical engineer who is designing the circuit, the specific design requirements for the MPDU were determined and used as guidelines. Space is limited due to the size of the mounting plate therefore each component must be strategically placed. Since the MPDU is being designed to fit into a simulated model of the spacecraft systems on the JIMO, components must be positioned where they are easily accessible to be wired to the other onboard systems. Mechanical and electrical requirements provided equally important limits which are combined to produce the best possible design of the MPDU.

Papa, Melissa R.↗
NASA NTRS2017

Robust Exploration and Commercial Missions to the Moon Using NTR LANTR Propulsion and Lunar-Derived Propellants

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. In his post-Apollo Integrated Space Program Plan (1970-1990), Wernher von Braun, proposed a reusable nuclear thermal propulsion stage (NTPS) to deliver cargo and crew to the Moon to establish a lunar base before undertaking human missions to Mars. The NTR option was selected by von Braun because it was a demonstrated technology capable of generating both high thrust and high specific impulse (Isp 900 s) twice that of todays best chemical rockets. In NASAs Mars Design Reference Architecture (DRA) 5.0 study, the crewed Mars transfer vehicle used three 25 klbf Pewee engines the smallest and highest performing engine tested in the Rover program along with graphite composite fuel. Smaller, lunar transfer vehicles consisting of a NTPS using three approximately 16.5 klbf Small Nuclear Rocket Engines (SNREs), an in-line propellant tank, plus the payload can enable a variety of reusable lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong tourism missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing an affordable in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The utilization of iron-rich volcanic glass or lunar polar ice (LPI) deposits (each estimated at billions of metric tons) for propellant production can significantly reduce the launch mass requirements from Earth and can enable reusable, surface-based lunar landing vehicles (LLVs) using liquid oxygen/hydrogen (LOX/LH2) chemical rocket engines. Afterwards, LOX/LH2 propellant depots can be established in lunar equatorial and polar orbits to supply the LTS. At this point a modified version of the conventional NTR called the LOX-augmented NTR, or LANTR would be introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an afterburner into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engines choked sonic throat essentially scramjet propulsion in reverse. By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and Isp values while the reactor core power level remains relatively constant. Eventually, a LANTR-based LTS can enable a rapid commuter shuttle with one-way trip times to and from the Moon ranging from 36 to 24 hours. Even if only 1 of the extracted propellant from identified volcanic glass and polar ice deposits were available for use in lunar orbit, such a supply could support daily commuter flights to the Moon for many thousands of years! An evolutionary mission architecture is outlined and a variety of lunar missions and transfer vehicle designs are examined, along with the increasing demands on propellant production as mission complexity increases. A comparison of vehicle features and engine operating characteristics, for both NTR and LANTR engines, is also provided along with a brief discussion on the propellant production issues associated with using volcanic glass and LPI as source material.

Spacecraft design↗
NASA NTRSDate not supplied

Hot Hydrogen Testing of W-coated dUN Kernels for Nuclear Thermal Propulsion

Tungsten-coated kernels of uranium nitride (UN) are a possible reactor fuel component for nuclear thermal propulsion. Hot hydrogen testing of such kernels (approx. 250μm diameter, coating thickness 5μm) has been performed at temperatures from 1800°C to 2300°C for 30min each, using depleted uranium nitride (dUN). In addition to direct observations and mass loss measurements, the samples were analyzed by XRD and SEM/EDS after each run. Decomposition of dUN and the formation of molten uranium happened at all temperatures tested with reaction rates increasing with temperature, despite the tungsten coating. At 2300°C, the kernels disintegrated and molten uranium seeped through the wall and bottom of the tungsten crucible employed. These hot hydrogen tests allowed to test the integrity of the tungsten coating and addressed the operating limits of the loose tungsten-coated dUN kernels. Improved results are expected upon consolidation of the tungsten-coated dUN kernels in a Mo-W alloy matrix.

nuclear fuels↗
NASA NTRSDate not supplied

In-Situ Resource Utilization Modeling of a Lunar Water Processing System

A key element of achieving a sustained surface presence, such as defined in NASA’s Artemis plan, is In-Situ Resource Utilization (ISRU). ISRU is the practice of using local resources to provide mission consumables that reduce system launch mass requirements, and regenerate resources (chiefly, water and oxygen) for propulsion and life support supporting both Lunar and Martian missions. ISRU systems require multiple complex processes, such as excavation, chemical reactors, and electrolysis subsystems that must operate in harmony to optimize the overall system process from beginning to end. The Mission Analysis and Integration Tool (MAIT) was previously developed with MATLAB in FY22 to connect individual subsystem models into a customized, flexible framework for the purpose of technology downselect, optimization, and end-to-end process planning. Beginning in FY24, MAIT was leveraged and evolved using MATLAB/Simulink due to its ability to communicate with a vast number of other programming languages and makes up the backbone of data flow between inputs and outputs to the subsystem models. MAIT initially evaluated a suite of ISRU-related technologies, including the water processing Lunar Auger Dryer for ISRU (LADI) system with integrated upstream excavation and downstream electrolysis subsystems. With individual models consolidated, the MAIT tool generated over 60,000 cases during its parametric sweeps; these system iterations produced valuable insight into the optimal LADI geometry for minimizing energy demands, estimating carbothermal reactor and radiator mass relationships, and calculated the power dynamics of the electrolysis unit. Advanced efforts with advanced models will include examining multiple production targets to demonstrate the ability to scale ISRU technologies supporting the Space Technology Mission Directorate’s (STMD) commercialization strategy, and increase the MAIT software capability to handle a wide array of ISRU system models beyond the Lunar environment, e.g. production of propellant for a Martian lander.

Avery Carlson↗
NASA NTRS2018

Robust Exploration and Commercial Missions to the Moon Using Nuclear Thermal Rocket Propulsion and Lunar Liquid Oxygen Derived from FeO-Rich Pyroclasitc Deposits

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. It is a demonstrated technology capable of generating both high thrust and high specific impulse (I(sub sp) approx. 900 s) twice that of today's best chemical rockets. Nuclear lunar transfer vehicles-consisting of a propulsion stage using three approx. 16.5-klb(sub f) small nuclear rocket engines (SNREs), an in-line propellant tank, plus the payload-are reusable, enabling a variety of lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong ''tourism'' missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing a robust in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The use of lunar liquid oxygen (LLO2) derived from iron oxide (FeO)-rich volcanic glass beads, found in numerous pyroclastic deposits on the Moon, can significantly reduce the launch mass requirements from Earth by enabling reusable, surface-based lunar landing vehicles (LLVs)that use liquid oxygen and hydrogen (LO2/LH2) chemical rocket engines. Afterwards, a LO2/LH2 propellant depot can be established in lunar equatorial orbit to supply the LTS. At this point a modified version of the conventional NTR-called the LO2-augmented NTR, or LANTR-is introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an ''afterburner'' into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engine's choked sonic throat-essentially ''scramjet propulsion in reverse.'' By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and I(sub sp) values while the reactor core power level remains relatively constant. A LANTR-based LTS offers unique mission capabilities including short-transit-time crewed cargo transports. Even a ''commuter'' shuttle service may be possible allowing ''one-way'' trip times to and from the Moon on the order of 36 hours or less. If only 1% of the extracted LLO2 propellant from identified resource sites were available for use in lunar orbit, such a supply could support daily commuter flights to the Moon for many thousands of years! This report outlines an evolutionary architecture and examines a variety of mission types and transfer vehicle designs, along with the increasing demands on LLO2 production as mission complexity and velocity change delta V requirements increase. A comparison of vehicle features and engine operating characteristics, for both NTR and LANTR engines, is also provided along with a discussion of the propellant production and mining requirements associated with using FeO-rich volcanic glass as source material.

rocket↗
NASA NTRS2002

Fusion for Space Propulsion

The need for fusion propulsion for interplanetary flights is discussed. For a propulsion system, there are three important system attributes: (1) The absolute amount of energy available, (2) the propellant exhaust velocity, and (3) the jet power per unit mass of the propulsion system (specific power). For efficient and affordable human exploration of the solar system, propellant exhaust velocity in excess of 100 km/s and specific power in excess of 10 kW/kg are required. Chemical combustion obviously cannot meet the requirement in propellant exhaust velocity. Nuclear fission processes typically result in producing energy in the form of heat that needs to be manipulated at temperatures limited by materials to about 2,800 K. Using the fission energy to heat a low atomic weight propellant produces propellant velocity of the order of 10 kinds. Alternatively the fission energy can be converted into electricity that is used to accelerate particles to high exhaust velocity. However, the necessary power conversion and conditioning equipment greatly increases the mass of the propulsion system. Fundamental considerations in waste heat rejection and power conditioning in a fission electric propulsion system place a limit on its jet specific power to the order of about 0.2 kW/kg. If fusion can be developed for propulsion, it appears to have the best of all worlds - it can provide the largest absolute amount of energy, the propellant exhaust velocity (> 100 km/s), and the high specific jet power (> 10 kW/kg). An intermediate step towards fusion propulsion might be a bimodal system in which a fission reactor is used to provide some of the energy to drive a fusion propulsion unit. There are similarities as well as differences between applying fusion to propulsion and to terrestrial electrical power generation. The similarities are the underlying plasma and fusion physics, the enabling component technologies, the computational and the diagnostics capabilities. These physics and engineering capabilities have been demonstrated for a fusion reactor gain (Q) of the order of unity (TFTR: 0.25, JET: 0.65, JT-60: Q(sub eq) approx. 1.25). These technological advances made it compelling for considering fusion for propulsion.

Thio, Y. C. Francis↗
NASA NTRS2018

Robust Exploration and Commercial Missions to the Moon Using Nuclear Thermal Rocket Propulsion and In Situ Propellants Derived from Lunar Polar Ice Deposits

The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable “access through cislunar space” necessary for commercial development and sustained human presence on the Moon. It is a demonstrated technology capable of generating both high thrust and high specific impulse (I(sub sp) ~900 s)— twice that of today’s best chemical rockets. Nuclear lunar transfer vehicles—consisting of a propulsion stage using three ~16.5-klb(sub f) small nuclear rocket engines (SNREs), an in-line propellant tank, plus the payload—can enable a variety of reusable lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong “tourism” missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing a robust in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in situ propellant production used to supply strategically located propellant depots and transportation nodes. The processing of lunar polar ice (LPI) deposits (estimated to be ~2 billion metric tons) for propellant production—specifically liquid oxygen (LO(sub 2)) and hydrogen (LH(sub 2))—can significantly reduce the launch mass requirements from Earth and can enable reusable, surface-based lunar landing vehicles (LLVs) using LO(sub 2)/LH(sub 2) chemical rocket engines. Afterwards, LO(sub 2)/LH(sub 2) propellant depots can be established in lunar polar and equatorial orbits to supply the LTS. At this point a modified version of the conventional NTR called the LO(sub 2)-augmented NTR, or LANTR, would be introduced into the LTS, allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants (LDPs) for Earth return. The bipropellant LANTR engine utilizes the large divergent section of its nozzle as an “afterburner” into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engine’s choked sonic throat—essentially “scramjet propulsion in reverse.” By varying the oxygen-to-hydrogen mixture ratio, LANTR engines can operate over a range of thrust and I(sub sp) values while the reactor core power level remains relatively constant. A LANTR-based LTS offers unique mission capabilities including short transit time crewed cargo transports. Even a “commuter” shuttle service may be possible, allowing “one-way” trip times to and from the Moon on the order of 36 hr or less. If only 1% of the postulated trapped water ice were available for use in lunar orbit, such a supply could support routine commuter flights to the Moon for many thousands of years. This report outlines an evolving LTS architecture that uses propellants derived from LPI and examines a variety of mission types and transfer vehicle designs along with their operating characteristics and increasing demands on LDP production as mission complexity and velocity change ΔV requirements increase. A comparison of the LDP production and mining requirements using LPI and volcanic glass to produce lunar-derived liquid oxygen (LUNOX) via the hydrogen reduction process is included, and the synergy with an evolving helium-3 mining industry is also discussed.

Borowski, Stanley K.↗
DOE OSTI2023

Nuclear fusion powered Titan aircraft

This paper discusses a system for Titan exploration enabled by nuclear fusion power. Titan is one of the most interesting locations in the solar system with a thick atmosphere, surface oceans, under-ice oceans and complex terrain. This paper provides a conceptual design of a fusion-powered system to explore many parts of Titan and enable the use of high-power instruments. The design includes a fusion-powered orbital transfer vehicle and an electric Titan science aircraft. A Direct Fusion Drive (DFD) propulsive engine could bring a sizable spacecraft to Titan orbit in less than two years. A second fusion reactor, configured as a closed-loop power generator, would be used for an electric Titan science aircraft. Both reactors are based on the Princeton Field-Reversed Configuration (PFRC) concept which combines an FRC with a magnetic mirror. PFRC uses a novel radio-frequency plasma heating system and deuterium-helium-3 fuel. A lower temperature plasma flows around the closed-field FRC region removing the fusion products. In the DFD propulsive configuration, this secondary flow permits direct and variable thrust and exhaust velocity. The science aircraft would do a powered entry to Titan and then have the capability to fly anywhere on the moon at subsonic speeds. The DFD-powered transfer vehicle would allow the in-orbit transfer stage to change inclination as needed to cover different areas of the surface.

33 ADVANCED PROPULSION SYSTEMS↗
NASA NTRS2019

The Value Proposition of Multi-Megawatt Electric Power/Propulsion for the Human Exploration of Mars

NASA and other space agencies have been offering stakeholders an architecture for the human exploration of Mars that has remained essentially unchanged since the 1960's. The mission duration and launch mass are the two first order figures of merit that drive the total cost of such an architecture. The options studied to date center on the "conjunction class" or "Long Stay" mission, in which surface infrastructure elements are sent ahead on uncrewed, slow trajectories requiring a minimum amount of propellant, and the crewed elements are sent at the synotic cycle's shortest trajectory, stay on Mars until the next close alignment, and then return to Earth. Total crewed mission durations for these architectures range from 900 to 1100 days, with variations driven primarily by trades between the amount of propellant launched and the effective specific mass of the propulsion technology assumed (e.g., chemical, nuclear thermal, solar electric, nuclear electric). Mars transit propulsion systems assumed in mission architectures studied to date have all resulted in architectural figures of merit that drive the cost to a level of "too much." However, nuclear electric propulsion (NEP) technology offers a "knob" that might be turned to enable a radically different Mars architecture, whose launch mass and mission duration may enable a value proposition more palatable to mission stakeholders. Mars architectures studied to date have assumed an NEP system providing 2.5 MWe at a specific mass of no less than 20 kg/kWe. This has often been seen as obtainable with a moderately high temperature fission reactor. An NEP system providing 15 MWe at specific mass of ~1 kg/kWe, though, could enable a short stay (30 days on surface) Mars mission, requiring only two or three SLS-class launch vehicles and a total mission duration of under one calendar year. However, turning this NEP "knob" would require a high risk development program driving innovation on the order of that delivered by the Manhattan Project, but for a fraction of the cost. Any technology option that might offer such a capability at such a development cost now stands at a low Technology Readiness Level (TRL) 3, would be based on a nuclear energy source (likely fusion), and would require an extremely high risk (and rapid) development effort. The aggressive, parallel path project management paradigm exemplified by the original Manhattan Project might have the best chance of success. An energy source developed in this manner may also have a major impact on the terrestrial power industry.

power↗
NASA NTRSDate not supplied

In-Situ Resource Utilization Modeling of a Lunar Water Processing System

A key element of achieving a sustained surface presence, such as defined in NASA’s Artemis plan, is In-Situ Resource Utilization (ISRU). ISRU is the practice of using local resources to provide mission consumables that reduce system launch mass requirements, and regenerate resources (chiefly, water and oxygen) for propulsion and life support supporting both Lunar and Martian missions. ISRU systems require multiple complex processes, such as excavation, chemical reactors, and electrolysis subsystems that must operate in harmony to optimize the overall system process from beginning to end. The Mission Analysis and Integration Tool (MAIT) was previously developed with MATLAB in FY22 to connect individual subsystem models into a customized, flexible framework for the purpose of technology downselect, optimization, and end-to-end process planning. Beginning in FY24, MAIT was updated and became the capital program in the Systems Engineering and Integration (SE&I) ISRU Modeling and Analysis (SIMA) project. Prior work was leveraged and evolved using MATLAB/Simulink due to its ability to communicate with a vast number of other programming languages and makes up the backbone of data flow between inputs and outputs to the subsystem models. MAIT initially evaluated a suite of ISRU-related technologies, including the water processing Lunar Auger Dryer for ISRU (LADI) system with integrated upstream excavation and downstream electrolysis subsystems. With individual models consolidated, the MAIT tool generated over 5,000 cases during its first round of parametric sweeps on the water processing architecture at multiple production targets; the system analysis produced valuable insight into the optimal LADI geometry that minimized energy demands, estimated effects to cold trap size and radiator requirements, and calculated the power dynamics of the electrolysis unit and liquid oxygen storage volume. Additional efforts are being made to demonstrate the ability to scale ISRU technologies supporting the Space Technology Mission Directorate’s (STMD) commercialization strategy and increase the MAIT software capability. Work is ongoing to handle a wide array of ISRU system models beyond the Lunar environment, e.g. production of propellant for a Martian lander.

Avery Carlson↗
NASA NTRSDate not supplied

In-Situ Resource Utilization Modeling of a Lunar Water Processing System

A key element of achieving a sustained surface presence, such as defined in NASA’s Artemis plan, is In-Situ Resource Utilization (ISRU). ISRU is the practice of using local resources to provide mission consumables that reduce system launch mass requirements, and regenerate resources (chiefly, water and oxygen) for propulsion and life support supporting both Lunar and Martian missions. ISRU systems require multiple complex processes, such as excavation, chemical reactors, and electrolysis subsystems that must operate in harmony to optimize the overall system process from beginning to end. The Mission Analysis and Integration Tool (MAIT) was previously developed with MATLAB in FY22 to connect individual subsystem models into a customized, flexible framework for the purpose of technology downselect, optimization, and end-to-end process planning. Beginning in FY24, MAIT was updated and became the capital program in the Systems Engineering and Integration (SE&I) ISRU Modeling and Analysis (SIMA) project. Prior work was leveraged and evolved using MATLAB/Simulink due to its ability to communicate with a vast number of other programming languages and makes up the backbone of data flow between inputs and outputs to the subsystem models. MAIT initially evaluated a suite of ISRU-related technologies, including the water processing Lunar Auger Dryer for ISRU (LADI) system with integrated upstream excavation and downstream electrolysis subsystems. With individual models consolidated, the MAIT tool generated over 5,000 cases during its first round of parametric sweeps on the water processing architecture at multiple production targets; the system analysis produced valuable insight into the optimal LADI geometry that minimized energy demands, estimated effects to cold trap size and radiator requirements, and calculated the power dynamics of the electrolysis unit and liquid oxygen storage volume. Additional efforts are being made to demonstrate the ability to scale ISRU technologies supporting the Space Technology Mission Directorate’s (STMD) commercialization strategy and increase the MAIT software capability. Work is ongoing to handle a wide array of ISRU system models beyond the Lunar environment, e.g. production of propellant for a Martian lander.

Avery Carlson↗
NASA NTRS2012

Near Earth Asteroid Human Mission Possibilities Using Nuclear Thermal Rocket (NTR) Propulsion

The NTR is a proven technology that generates high thrust and has a specific impulse (Isp (is) approximately 900 s) twice that of today's best chemical rockets. During the Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) programs, twenty rocket reactors were designed, built and ground tested. These tests demonstrated: (1) a wide range of thrust; (2) high temperature carbide-based nuclear fuel; (3) sustained engine operation; (4) accumulated lifetime; and (5) restart capability - all the requirements needed for a human mission to Mars. Ceramic metal fuel was also evaluated as a backup option. In NASA's recent Mars Design reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, versatile vehicle design, simple assembly, and growth potential. In contrast to other advanced propulsion options, NTP requires no large technology scale-ups. In fact, the smallest engine tested during the Rover program - the 25 klbf 'Pewee' engine is sufficient for a human Mars mission when used in a clustered engine configuration. The 'Copernicus crewed NTR Mars transfer vehicle design developed for DRA 5.0 has significant capability that can enable reusable '1-year' round trip human missions to candidate near Earth asteroids (NEAs) like 1991 JW in 2027, or 2000 SG344 and Apophis in 2028. A robotic precursor mission to 2000 SG344 in late 2023 could provide an attractive Flight Technology Demonstration of a small NTR engine that is scalable to the 25 klbf-class engine used for human missions 5 years later. In addition to the detailed scientific data gathered from on-site inspection, human NEA missions would also provide a valuable 'check out' function for key elements of the NTR transfer vehicle (its propulsion module, TransHab and life support systems, etc.) in a 'deep space' environment prior to undertaking the longer duration Mars orbital and landing missions that would follow. The initial mass in low Earth orbit required for a mission to Apophis is approximately 323 t consisting of the NTR propulsion module ((is) approximately 138 t), the integrated saddle truss and LH2 drop tank assembly ((is) approximately 123 t), and the 6-crew payload element ((is) approximately 62 t). The later includes a multi-mission Space Excursion Vehicle (MMSEV) used for close-up examination and sample gathering. The total burn time and required restarts on the three 25 klbf 'Pewee-class' engines operating at Isp (is) approximately 906 s, are approximately 76.2 minutes and 4, respectively, well below the 2 hours and 27 restarts demonstrated on the NERVA eXperimental Engine, the NRX-XE. The paper examines the benefits, requirements and characteristics of using NTP for the above NEA missions. The impacts on vehicle design of HLV payload volume and lift capability, crew size, and reusability are also quantified.

Borowski, Stanley↗