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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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All sourcesNASA NTRSDOE OSTI
DOE OSTI2025

Modeling a Sodium Heat Pipe Experiment at SPHERE Using Sockeye

The Single Primary Heat Extraction and Rejection Emulator (SPHERE) facility at Idaho Na- tional Laboratory was recently utilized to generate data for the startup and steady operation of a high-performance, sodium heat pipe over the course of 1000 hours, as a test of detrimental, long-term effects of heat pipe operation. The setup consists of a single, sodium heat pipe enclosed in a stainless-steel vacuum chamber, heated radiatively via a cylindrical ceramic fiber heater configuration and cooled via a water-cooled calorimeter. Measurements include temperatures at several axial locations along the outer surface of the heat pipe, the power provided to the heaters, and the heat removal rate of the calorimeter. In this work, this data is utilized to validate heat pipe models in the heat pipe application Sockeye, which is based upon the Multiphysics Object-Oriented Simulation Environment (MOOSE) framework. Sockeye provides various heat pipe models at an engineering scale appropriate for the multiphysics simulation of microreactors, which may feature several hundred heat pipes. This work details models of this experiment in SPHERE using various heat pipe models with Sockeye, including heat conduction-based models and compressible flow models of the heat pipe interior. These models are compared to the experimental data to assess the accuracy of several aspects of heat pipe modeling, including frozen startup, the effect of non-condensable gases, and the coupling of the heat pipe to its environment.

97 - MATHEMATICS AND COMPUTING↗
NASA NTRS2018

Creating a Voyager Thermal Model 39 Years Into the Flight Mission, Along with Model Correlation and Application

After 39 years of continuous operation in space, the output of the Voyager 1 & 2 spacecraft RTG power systems has decreased to the point where managing the power margin and maintaining thermal control has become increasingly difficult. As the total power dissipation in the bus has decreased, propellant line temperatures and margin above minimum AFTs have decreased, creating risk of the hydrazine freezing (at 1.6°C). This is further complicated by the lack of existing thermal models that can be used to assess propellant tank and line temperatures. In 2014, an effort was begun to create a Voyager spacecraft thermal model for that purpose.A steady-state Thermal Desktop model has been created from scratch over the past two years. The initial thermal model development was started by Applied Sciences Laboratory (ASL) under contract to JPL. The effort relied primarily on archived manufacturing drawings, limited documentation, interviews of senior engineers who worked on the Voyager design and implementation, and the experience of the Voyager Flight Operations team.Data from the Voyager System Thermal Vacuum tests is no longer available, making it necessary to correlate the model to more recent flight data and small in-flight tests. Correlation was achieved to within ±5°C for a hot case and a cold case (both data sets from 2014). However, the flight system has very few temperature sensors directly on propellant lines. So the task remains to determine how best to use the model, in conjunction with flight data, to make sure the Voyagers can continue to fly successfully. How does one go about creating a thermal model for a spacecraft that is already launched, has limited existing mechanical description files, no thermal model in order to do a maneuver that was never planned when the spacecraft was designed?

Ledeboer, William C.↗
NASA NTRS2016

Advanced Stirling Radioisotope Generator EU2 Anomaly Investigation

The Advanced Stirling Radioisotope Generator (ASRG) Engineering Unit 2 (EU2) is the highest fidelity electrically-heated Stirling radioisotope generator built to date. NASA Glenn Research Center (GRC) completed the assembly of the ASRG EU2 in September, 2014 using hardware from the now cancelled ASRG flight development project. The ASRG EU2 integrated the first pair of Sunpower's ASC-E3 Stirling convertors (ASC-E3 #1 and #2) in an aluminum generator housing with Lockheed Martin's Engineering Development Unit (EDU) 4 controller. After just 179 hours of EU2 generator operation, the first power fluctuation occurred on ASC-E3 #1. The first power fluctuation occurred 175 hours later on ASC-E3 #2. Over time, the power fluctuations became more frequent on both convertors and larger in magnitude. Eventually the EU2 was shut down in January, 2015. An anomaly investigation was chartered to determine root cause of the power fluctuations and other anomalous observations. A team with members from GRC, Sunpower, and Lockheed Martin conducted a thorough investigation of the EU2 anomalies. Findings from the EU2 disassembly identified proximate causes of the anomalous observations. Discussion of the team's assessment of the primary possible failure theories, root cause, and conclusions is provided. Recommendations are made for future Stirling generator development to address the findings from the anomaly investigation. Additional findings from the investigation are also discussed.

Lewandowski, Edward J.↗
NASA NTRS2018

Optically Inaccessible Flow Visualization using Positron Emission Tomography

Flow visualization is a powerful tool for characterizing fluid dynamics within engineering systems that utilize fluid working media. Recent advances in Positron Emission Tomography (PET) have enhanced its ability to extend beyond the medical field, and offer an alternate vantage point in visualizing optically inaccessible fluid distributions and flow fields within the aerospace field. In light of this prospect an investigation has ensued to parametrically bound the flows that can be sufficiently resolved using current PET technology. Results from an initial series of experiments and analyses performed at the Positron Imaging Centre (PIC) located at the University of Birmingham, UK, will be presented. Discussion will also cover preliminary, on-going, activities associated with assessing the utility of PET technology in zero-gravity propellant gauging, ionization efficiency characterization in electric propulsion devices, and broader industrial applications.

PET↗
DOE OSTI1992

Nuclear Safety [Vol. 33, No. 4, October-December 1992]

Nuclear Safety is a review journal that covers significant developments in the field of nuclear safety. Its scope includes the analysis and control of hazards associated with nuclear energy, operations involving fissionable materials, and the products of nuclear fission and their effects on the environment. Primary emphasis is on safety in reactor design, construction, and operation; however, the safety aspects of the entire fuel cycle, including fuel fabrication, spent-fuel processing, nuclear waste disposal, handling of radioisotopes, and environmental effects of these operations, are also treated. Table of Contents for this issue follows. GENERAL SAFETY CONSIDERATIONS: 499 Technical Note: Nuclear Regulatory Commission’s Examination Question Bank System, C. L Jensen and N. L. Skinner; 506 Report on the Specialists’ Meeting on Passive and Active Safety Features of Liquid-Metal Fast Breeder Reactors Organized by the International Atomic Energy Agency at Oarai Engineering Centre of Power Reactor and Nuclear Development Corporation, Japan, November 5-7,1991, S. R. Paranjpe; ACCIDENT ANALYSIS: 514 A Review of Hydrogen Production During Melt/Water Interaction in LWRs, D. F. Fletcher, B. D. Turland, and S. P. A. Lawrence; 535 A Review of the U 3 0 8 -Al Reaction as a Potential Heat Source in Research and Test Reactor Accidents, J. L. Snelgrove; CONTROL AND INSTRUMENTATION: 543 Multilevel Interfaces for Power Plant Control Rooms II: A Preliminary Design Space, K. J. Vicente; DESIGN FEATURES: 549 Twenty-Second DOE/NRC Nuclear Air Cleaning and Treatment Conference, R. R. Bellamy, D. W. Moeller, and M. W. First; 562 Book Review: Safety Assessment of Unit 5 (WWER-440/W-213) of the Greifswald Nuclear Power Station, M. H. Fontana; WASTE AND SPENT FUEL MANAGEMENT: 576 Activities Related to Waste and Spent Fuel Management, M. D. Muhlheim and E. G. Silver; OPERATING EXPERIENCES: 586 Analysis of Corrective Actions Applied to Nuclear Power Plant Operations, A. E. Cross; 593 Reactor Shutdown Experience, Compiled by J. W. Cletcher; 596 Operating U.S. Power Reactors, Compiled by M. D. Muhlheim and E. G. Silver; 609 Non-Commercial Power Operating Experience; 614 World Nuclear Performance, Compiled by M. D. Muhlheim; RECENT DEVELOPMENTS: 623 General Administrative Activities, Compiled by M. D. Muhlheim and E. G. Silver; 635 Reports, Standards, and Safety Guides, D. S. Queener; 639 Proposed Rule Changes as of June 30, 1992; ANNOUNCEMENTS: 513 Fourth International Conference on Nuclear Fuel Reprocessing and Waste Management RECOD '94; 585 International Topical Meeting on Advanced Reactors Safety; 622 International Workshop on Implementation of ALARA at Nuclear Power Plants; 648 Short Course on Nuclear Criticality Safety; 648 8th Topical Meeting on Problems in Reactor Physics MEPhi-VOLGA-93; 643 The Authors; 646 Reviewers of Nuclear Safety, Vol. 33

11 NUCLEAR FUEL CYCLE AND FUEL MATERIALS↗
NASA NTRS1979

A turbojet-boosted two-stage-to-orbit space transportation system design study

The concept to use twin turbo-powered boosters for acceleration to supersonic staging speed followed by an all rocket powered orbiter stage was proposed. A follow-on design study was then made of the concept with the performance objective of placing a 29,483 Kg payload into a .2.6 X 195.3 km orbit. The study was performed in terms of analysis and trade studies, conceptual design, utility and economic analysis, and technology assessment. Design features of the final configuration included: strakes and area rule for improved take off and low transonic drag, variable area inlets, exits and turbine, and low profile fixed landing gear for turbojet booster stage. The payload required an estimated GLOW of 1,270,000 kg for injection in orbit. Each twin booster required afterburning turbojet engines each with a static sea level thrust rating of 444,800 N. Life cycle costs for this concept were comparable to a SSTO/SLED concept except for increased development cost due to the turbojet engine propulsion system.

Hepler, A. K.↗
NASA NTRS1982

Broad specification fuels technology program, phase 1

An experimental evaluation was conducted to assess the impact of the use of broadened properties fuels on combustor design concepts. Emphasis was placed on establishing the viability of design modifications to current combustor concepts and the use of advanced technology concepts to facilitate operation on Experimental Referee Broad Specification (ERBS) fuel while meeting exhaust emissions and performance specifications and maintaining acceptable durability. Three different combustor concepts, representative of progressively more aggressive technology levels, were evaluated. When operated on ERBS rather than Jet A fuel, a single stage combustor typical of that in the most recent versions of the JT9D-7 engine was found to produce excess carbon monoxide emissions at idle and elevated liner temperatures at high power levels that were projected to reduced liner life by 13 percent. The introduction of improved component technology, such as refined fuel injectors and advanced liner cooling concepts were shown to have the potential of enhancing the fuel flexibility of the single stage combustor.

Lohmann, R. P.↗
NASA NTRS1991

A quick accurate model of nozzle backflow

Backflow from nozzles is a major source of contamination on spacecraft. If the craft contains any exposed high voltages, the neutral density produced by the nozzles in the vicinity of the craft needs to be known in order to assess the possibility of Paschen breakdown or the probability of sheath ionization around a region of the craft that collects electrons for the plasma. A model for backflow has been developed for incorporation into the Environment-Power System Analysis Tool (EPSAT) which quickly estimates both the magnitude of the backflow and the species makeup of the flow. By combining the backflow model with the Simons (1972) model for continuum flow it is possible to quickly estimate the density of each species from a nozzle at any position in space. The model requires only a few physical parameters of the nozzle and the gas as inputs and is therefore ideal for engineering applications.

Kuharski, R. A.↗
NASA NTRS2007

Electric Propulsion Upper-Stage for Launch Vehicle Capability Enhancement

The NASA In-Space Propulsion Technology Project Office initiated a preliminary study to evaluate the performance benefits of a solar electric propulsion (SEP) upper-stage with existing and near-term small launch vehicles. The analysis included circular and elliptical Low Earth Orbit (LEO) to Geosynchronous Earth Orbit (GEO) transfers, and LEO to Low Lunar Orbit (LLO) applications. SEP subsystem options included state-of-the-art and near-term solar arrays and electric thrusters. In-depth evaluations of the Aerojet BPT-4000 Hall thruster and NEXT gridded ion engine were conducted to compare performance, cost and revenue potential. Preliminary results indicate that Hall thruster technology is favored for low-cost, low power SEP stages, while gridded-ion engines are favored for higher power SEP systems unfettered by transfer time constraints. A low-cost point design is presented that details one possible stage configuration and outlines system limitations, in particular fairing volume constraints. The results demonstrate mission enhancements to large and medium class launch vehicles, and mission enabling performance when SEP system upper stages are mounted to low-cost launchers such as the Minotaur and Falcon 1. Study results indicate the potential use of SEP upper stages to double GEO payload mass capability and to possibly enable launch on demand capability for GEO assets. Transition from government to commercial applications, with associated cost/benefit analysis, has also been assessed. The sensitivity of system performance to specific impulse, array power, thruster size, and component costs are also discussed.

Kemp, Gregory E.↗
DOE OSTI2025

On-target uniformity of the OMEGA 60-beam inertial confinement fusion laser

Successful direct-drive inertial confinement fusion (ICF) experiments require excellent on-target laser 9 irradiance uniformity maintained over the duration of the pulse. Achieving this symmetry relies on maintaining an 10 energy, power, and fluence balance among all beams in a multibeam laser system. As a result of the improvements 11 described in this paper, the OMEGA 60-beam laser performance has been assessed at approximately 2% on-target 12 irradiance nonuniformity using an updated performance assessment metric that accounts for the laser diagnostic noise 13 floor. The performance is measurement-limited, prompting the need for an improved diagnostic suite. This manuscript 14 provides a comprehensive first-order assessment of extant status of on-target energy, power, and fluence uniformity 15 and explores avenues for improvements.

Diagnostics↗
NASA NTRS2014

Mars Sample Return Using Commercial Capabilities: Mission Architecture Overview

Mars Sample Return (MSR) is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. This paper presents an overview of a feasibility study for an MSR mission. The objective of the study was to determine whether emerging commercial capabilities can be used to reduce the number of mission systems and launches required to return the samples, with the goal of reducing mission cost. We report the feasibility of a complete and closed MSR mission design using the following scenario that covers three synodic launch opportunities, beginning with the 2022 opportunity: A Falcon Heavy injects a SpaceX Red Dragon capsule and trunk onto a Trans Mars Injection (TMI) trajectory. The capsule is modified to carry all the hardware needed to return samples collected on Mars including a Mars Ascent Vehicle (MAV), an Earth Return Vehicle (ERV), and hardware to transfer a sample collected in a previously landed rover mission to the ERV. The Red Dragon descends to land on the surface of Mars using Super Sonic Retro Propulsion (SSRP). After previously collected samples are transferred to the ERV, the single-stage MAV launches the ERV from the surface of Mars. The MAV uses a storable liquid bi-propellant propulsion system to deliver the ERV to a Mars phasing orbit. After a brief phasing period, the ERV, which also uses a storable bi-propellant system, performs a Trans Earth Injection (TEI) burn. Upon arrival at Earth, the ERV performs Earth and lunar swing-bys and is placed into a lunar trailing circular orbit - an Earth orbit, at lunar distance. A later mission, using Dragon and launched by a Falcon Heavy, performs a rendezvous with the ERV in the lunar trailing orbit, retrieves the sample container and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft makes a controlled Earth re-entry preventing any unintended release of pristine martian materials into the Earth's biosphere. The analysis methods employed standard and specialized aerospace engineering tools. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships (MERs). The architecture was iterated until overall mission convergence was achieved on at least one path. Subsystems analyzed in this study include support structures, power system, nose fairing, thermal insulation, actuation devices, MAV exhaust venting, and GN&C. Best practice application of loads, mass growth contingencies, and resource margins were used. For Falcon Heavy capabilities and Dragon subsystems we utilized publically available data from SpaceX, published analyses from other sources, as well as our own engineering and aerodynamic estimates. Earth Launch mass is under 11 mt, which is within the estimated capability of a Falcon Heavy, with margin. Total entry masses between 7 and 10 mt were considered with closure occurring between 9 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance - have been derived. An assessment of the effect of the entry conditions on the thermal protection system (TPS), currently in use for Dragon missions, shows no significant stressors. A useful payload mass of 2.0 mt is provided and includes mass growth allowances for the MAV, the ERV, and mission unique equipment. We also report options for the MAV and ERV, including propulsion systems, crewed versus robotic retrieval mission, as well as direct Earth entry. International planetary protection policies as well as verifiable means of compliance will have a large impact on any MSR mission design. We identify areas within our architecture where such impacts occur. We also describe preliminary compliance measures that will be the subject of future work. This work shows that emerging commercial capabilities as well as new methodologies can be used to efficiently support an important planetary science objective. The work also has applications for human exploration missions that use propulsive EDL techniques

Red Dragon↗
NASA NTRS2013

Mars Sample Return Using Commercial Capabilities: Mission Architecture Overview

Mars Sample Return (MSR) is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. This paper presents an overview of a feasibility study for a MSR mission. The objective of the study was to determine whether emerging commercial capabilities can be used to reduce the number of mission systems and launches required to return the samples, with the goal of reducing mission cost. The major element required for the MSR mission are described and include an integration of the emerging commercial capabilities with small spacecraft design techniques; new utilizations of traditional aerospace technologies; and recent technological developments. We report the feasibility of a complete and closed MSR mission design using the following scenario that covers three synodic launch opportunities, beginning with the 2022 opportunity: A Falcon Heavy injects a SpaceX Red Dragon capsule and trunk onto a Trans Mars Injection (TMI) trajectory. The capsule is modified to carry all the hardware needed to return samples collected on Mars including a Mars Ascent Vehicle (MAV); an Earth Return Vehicle (ERV); and hardware to transfer a sample collected in a previously landed rover mission to the ERV. The Red Dragon descends to land on the surface of Mars using Supersonic Retro Propulsion (SRP). After previously collected samples are transferred to the ERV, the single-stage MAV launches the ERV from the surface of Mars to a Mars phasing orbit. The MAV uses a storable liquid, pump fed bi-propellant propulsion system. After a brief phasing period, the ERV, which also uses a storable bi-propellant system, performs a Trans Earth Injection (TEI) burn. Once near Earth the ERV performs Earth and lunar swing-bys and is placed into a Lunar Trailing Orbit (LTO0 - an Earth orbit, at lunar distance. A later mission, using a Dragon and launched by a Falcon Heavy, performs a rendezvous with the ERV in the lunar trailing orbit, retrieves the sample container and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft, makes a controlled Earth re-entry preventing any unintended release of pristine Martian materials into the Earth's biosphere. Other capsule type vehicles and associated launchers may be applicable. The analysis methods employed standard and specialized aerospace engineering tools. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships (MERs). The architecture was iterated until overall mission convergence was achieved on at least one path. Subsystems analyzed in this study include support structures, power system, nose fairing, thermal insulation, actuation devices, MAV exhaust venting, and GN&C. Best practice application of loads, mass growth contingencies, and resource margins were used. For Falcon Heavy capabilities and Dragon subsystems we utilized publically available data from SpaceX; published analyses from other sources; as well as our own engineering and aerodynamic estimates. Earth Launch mass is under 11 mt, which is within the estimated capability of a Falcon Heavy, with margin. Total entry masses between 7 and 10 mt were considered with closure occurring between 9 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance - have been derived. An assessment of the entry conditions on the thermal protection system (TPS), currently in use for Dragon missions, has been made. And shows no significant stressors. A useful mass of 2.0 mt is provided and includes mass growth allowances for the MAV, the ERV, and mission unique equipment. We also report on alternate propellant options for the MAV and options for the ERV, including propulsion systems; crewed versus robotic retrieval mission; as well as direct Earth entry. International Planetary Protection Policies as well as verifiable means of compliance will have a large impact on any MSR mission design. We identify areas within our architecture where such impacts occur. This work shows that emerging commercial capabilities can be used to effectively integrated into a mission to achieve an important planetary science objective.

Mars Sample Return↗
NASA NTRS2015

Electronegative Gas Thruster - Direct Thrust Measurement Project

This effort is an international collaboration and academic partnership to mature an innovative electric propulsion (EP) thruster concept to TRL 3 through direct thrust measurement. The initial target application is for Small Satellites, but can be extended to higher power. The Plasma propulsion with Electronegative GASES (PEGASES) concept simplifies ion thruster operation, eliminates a neutralizer requirement and should yield longer life capabilities and lower cost implementation over conventional gridded ion engines. The basic proof-of concept has been demonstrated and matured to TRL 2 over the past several years by researchers at the Laboratoire de Physique des Plasma in France. Due to the low maturity of the innovation, there are currently no domestic investments in electronegative gas thrusters anywhere within NASA, industry or academia. The end product of this Center Innovation Fund (CIF) project will be a validation of the proof-of-concept, maturation to TRL 3 and technology assessment report to summarize the potential for the PEGASES concept to supplant the incumbent technology. Information exchange with the foreign national will be one-way with the exception of the test results. Those test results will first go through a standard public release ITAR/export control review, and the results will be presented in a public technical forum, and the results will be presented in a public technical forum.

Dankanich, John↗
NASA NTRS2022

A Discussion of the Failure of a Quad Diode Module and Efforts to Assure the Flight Spares

In 2019 the International Space Station (ISS) experienced an on on-orbit failure that affected 1 of 28 Battery Charge Discharge Units (BCDUs). Telemetry pointed to a short circuit failure of a Power Rectifier Quad Diode Module. Astronauts removed the failed unit from service which was then returned to Earth for failure analysis. The failure analysis confirmed that the quad diode module had failed short circuit. The investigation identified silver dendrites had grown on the insulated, sloped edge of the mesa semiconductor die of the failed diode and also a 2nd diode in the same quad diode module. Voids between the diode’s protective encapsulating ring and the die provided space within which dendrites were able to form and cause catastrophic failure. The NASA Engineering & Safety Center (NESC) convened a EEE Parts Sub-Team to investigate root cause and to assist with risk assessment for all of the flight diode modules (4 distinct production lots) and the flight spares (from a 5th lot). Analysis of the original diode manufacturer’s read and record screening test data identified some lots which contained diodes with parametric instabilities that could signal conditions suited for silver dendrite formation. Computed Tomography (CT) X-ray screening inspection was performed on flight spare modules to identify diodes that do not exhibit voids between the edge of die and encapsulating ring as a mitigation against dendrite formation.

Jay Brusse↗
NASA NTRS1991

Navier-Stokes calculations of scramjet-nozzle-afterbody flowfields

A comprehensive computational fluid dynamics effort was conducted from 1987 to 1990 to properly design a nozzle and lower aft end of a generic hypersonic vehicle powered by a scramjet engine. The interference of the exhaust on the control surfaces of the vehicle can have adverse effects on its stability. Two-dimensional Navier-Stokes computations were performed, where the exhaust gas was assumed to be air behaving as a perfect gas. Then the exhaust was simulated by a mixture of Freon-12 and argon, which required solving the Navier-Stokes equations for four species: (nitrogen, oxygen, Freon-12, and argon). This allowed gamma to be a field variable during the mixing of the multispecies gases. Two different mixing models were used and comparisons between them as well as the perfect gas air calculations were made to assess their relative merits. Finally, the three dimensional Navier-Stokes computations were made for the full-span scramjet nozzle afterbody module.

Baysal, Oktay↗
NASA NTRS1996

Active Noise Control of Low Speed Fan Rotor-Stator Modes

This report describes the Active Noise Cancellation System designed by General Electric and tested in the NASA Lewis Research Center's 48 inch Active Noise Control Fan. The goal of this study was to assess the feasibility of using wall mounted secondary acoustic sources and sensors within the duct of a high bypass turbofan aircraft engine for active noise cancellation of fan tones. The control system is based on a modal control approach. A known acoustic mode propagating in the fan duct is cancelled using an array of flush-mounted compact sound sources. Controller inputs are signals from a shaft encoder and a microphone array which senses the residual acoustic mode in the duct. The canceling modal signal is generated by a modal controller. The key results are that the (6,0) mode was completely eliminated at 920 Hz and substantially reduced elsewhere. The total tone power was reduced 9.4 dB. Farfield 2BPF SPL reductions of 13 dB were obtained. The (4,0) and (4,1) modes were reduced simultaneously yielding a 15 dB modal PWL decrease. Global attenuation of PWL was obtained using an actuator and sensor system totally contained within the duct.

Sutliff, Daniel L.↗
DOE OSTI2025

Insulation Activation and Contamination in Low-Power Experiments

Among the diverse advanced nuclear reactor concepts being developed are designs that expect to operate at much higher temperatures than conventional light-water reactors, and therefore require alternative materials for components capable of withstanding these extreme temperatures. Insulation is one such component where alternative high-temperature industrial composites are being considered. Low-power experiments can provide an opportunity for advancing our understanding of material behavior under irradiation; however, these experiments are not without occupational hazards. The focus of this paper is the potential for activation of insulating composites when exposed to neutron fluence during reactor experiments, and to discuss the subsequent contamination of the testing area. This paper will also discuss practical measures for preventing the creation and inhalation of activated particulate matter (i.e., dust), such as engineering and administrative controls and personal protective equipment. This paper evaluates the suitability of several insulation materials for use in reactor experiments based on their activation from neutron irradiation. The materials considered are Pyrogel XTE by Aspen Aerogel, Cerablanket and Kaowool by Morgan Advanced Materials, Thermo-12 Gold by Johns Manville, and Maxsil CF6-2000 by McAllister Mills. Although these materials are not yet widely used in the nuclear industry, they are common in other industries due to their thermal insulating properties. These materials produce dust while handling that can pose health and safety hazards, which could be further exacerbated by activation. The analysis presented in this paper utilizes a deterministic method developed to predict the activation source term of the provided insulation materials. The method calculates the atomic density of each element within the composites and simulates their exposure to a specified neutron flux density representative of low power experimental conditions. During the irradiation phase and throughout the extensive cooling period, neutron absorptions and nuclear decays are tracked for relevant isotopes. The residual activity within each composite is then assessed to identify the material with the lowest potential radiological hazard to workers. Dust will be created during installation and removal of insulating materials and may also be generated during the experiment. Contamination of surrounding surfaces due to dispersion is highly likely unless preventative measures are taken. The results of this analysis are used to estimate an internal dose from inhalation of activated dust, which is significantly lower than the allowable whole-body dose. However, the risk of internal exposure may increase if the dust becomes a transport vector for nuclear material. The goals of this paper are to enhance the understanding of the potential activation and contamination hazards from the use of these insulating materials in low-power nuclear experiments and to foster safer work environments through informed decision-making and strategic planning. This work ultimately contributes to a broader understanding of how new non-nuclear materials for advanced reactors impact the radiological safety of co-located workers, helping to ensure the responsible management of radioactive materials during installation and decommissioning activities.

12 - MGMT OF RADIOACTIVE AND NON-RADIOACTIVE WASTE↗
NASA NTRS2018

Creating a Voyager Thermal Model 39 Years Into the Flight Mission, Along With Model Correlation and Application

After 39 years of continuous operation in space, the output of the Voyager 1 & 2 spacecraft Radioisotope Thermoelectric Generator (RTG) power systems has decreased to the point where managing the power margin and maintaining thermal control has become increasingly difficult. As the total power dissipation in the bus has decreased, propellant line temperatures and margin above minimum Allowable Flight Temperature (AFT) have decreased, creating risk of the hydrazine freezing (at 1.6°C). This is further complicated by the lack of existing thermal models that can be used to assess propellant tank and line temperatures. In 2014, an effort was begun to create a Voyager spacecraft thermal model for that purpose. A steady-state Thermal Desktop model has been created from scratch over the past two years. Applied Sciences Laboratory (ASL) started the initial thermal model development under contract to Jet Propulsion Laboratory (JPL). The effort relied primarily on archived manufacturing drawings, limited documentation, interviews of senior engineers who worked on the Voyager design and implementation, and the experience of the Voyager Flight Operations team. Data from the Voyager System Thermal Vacuum (STV) tests is no longer available, making it necessary to correlate the model to more recent flight data and small in-flight tests. Correlation was achieved to within ±5°C for a hot case and a cold case (both data sets from 2014). However, the flight system has very few temperature sensors directly on propellant lines. Therefore, the task remains to determine how best to use the model, in conjunction with flight data, to make sure the Voyagers can continue to fly successfully.

Ledeboer, William C.↗