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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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At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Comparison of Exploration Oxygen Recovery Technology Options Using ESM and LSMAC

In preparation for long duration manned space flight, numerous technology development efforts are ongoing in the area of environmental control and life support (ECLS). In cooperation with international, industry, and academic partners, NASA seeks to leverage the International Space Station as a testbed for technologies targeted for Exploration-class missions. In recent years, Equivalent Systems Mass (ESM) analyses have been conducted to evaluate the relative breakeven points and to compare technologies as part of ECLS architectural trades. While these studies have provided important data pertaining to key engineering metrics, additional considerations are important to more fully understand the potential impacts and costs associated with selecting a specific architecture. A tool, called the Life Support Multi-Dimensional Assessment Criteria (LSMAC), was recently proposed by Sierra Nevada Corporation in an attempt to incorporate influences of these additional considerations including Maintainability, Risk Analysis, Technology Readiness Level, Radiation Impacts, Manufacturing Costs, Reliability, Human Factors, and Un-Crewed Operations. As a first step toward evaluating and implementing this tool, LSMAC was used to revisit the ISS oxygen recovery trade from the 1990’s wherein Sabatier was selected over Bosch technology. Second, the tool was used to compare oxygen recovery developmental technologies currently in work. The results of these studies as well as a comparison with standalone ESM analyses are reported. Further, a discussion of the potential application of the tool across the ECLS portfolio and its potential use in future technology selection for ISS flight demonstrations is provided.

Morgan B Abney↗
NASA NTRSDate not supplied

Metrics in Space Life Support Technology Selection

Engineering metrics are useful in space life support technology selection, but they must be carefully used. Metrics are only part of a complete system trade-off. Metrics do harm if they cause neglect of other important technical, organizational, or intuitive decision factors. Two metrics have damaged space life support, closure and Equivalent Systems Mass (ESM). Closure measures the fraction of the required system inputs that are produced by recycling system outputs. Increasing closure produces diminishing returns and becomes increasingly expensive. Increasing closure does not directly contribute to providing better life support. ESM measures the total launch mass required to provide life support. ESM includes the mass of the system hardware and of its power, cooling, pressurized volume, spares, and logistics. ESM predicts launch costs, but recently launch costs have been reduced by a factor of 20 or more. System development cost for space hardware is often much greater than launch cost. The past nearly exclusive use of ESM has led to the neglect of Life Cycle Cost (LCC), reliability, cost, and the other engineering factors. Closure and ESM have misguided space life support technology selection for more than twenty years and have adversely affected the expenditure of 100’s of millions of dollars. Metrics can be effectively used three ways in space life support technology selection: 1. A small set of key engineering metrics for preliminary screening. 2. A full set of engineering to guide technical selection. 3. Combining engineering metrics with organizational, political, and intuitive decision factors to understand technology selection. The past emphasis on closure and ESM served to support recycling life support over resupply and built on the intuitive appeal of a human ecosystem in space.

Harry W Jones↗
NASA NTRSDate not supplied

The Effective Use of Metrics in Space Life Support System Trade-Offs

Engineering metrics are useful in space life support technology selection, but they must be carefully used. Metrics are only part of a complete system trade-off. Metrics do harm if they cause neglect of other important technical, organizational, or intuitive decision factors. Two metrics have damaged space life support, closure and Equivalent Systems Mass (ESM). Closure measures the fraction of the required system inputs that are produced by recycling system outputs. Increasing closure produces diminishing returns and becomes increasingly expensive. Increasing closure does not directly contribute to providing better life support. ESM measures the total launch mass required to provide life support. ESM includes the mass of the system hardware and of its power, cooling, pressurized volume, spares, and logistics. ESM predicts launch costs, but recently launch costs have been reduced by a factor of 20 or more. System development cost for space hardware is often much greater than launch cost. The past nearly exclusive use of ESM has led to the neglect of Life Cycle Cost (LCC), reliability, cost, and the other engineering factors. Closure and ESM have misguided space life support technology selection for more than twenty years and have adversely affected the expenditure of 100’s of millions of dollars. Metrics can be effectively used three ways in space life support technology selection: 1. A small set of key engineering metrics for preliminary screening. 2. A full set of engineering to guide technical selection. 3. Combining engineering metrics with organizational, political, and intuitive decision factors to understand technology selection. The past emphasis on closure and ESM served to support recycling life support over resupply and built on the intuitive appeal of a human ecosystem in space.

Harry W. Jones↗
NASA NTRSDate not supplied

Trade Study Analysis of a Cryogenic Oxygen Architecture for Lunar Outpost Life Support

A trade study was performed to compare the use of cryogenic liquid oxygen (LOX) with high pressure gaseous oxygen (GOX) and electrolysis approaches for Lunar outpost life support, which consists of a surface habitat and pressurized rover. This study presents the mission details pertaining to a Lunar outpost architecture, discusses the viable concept of operations for each architecture, and compares the equivalent system mass (ESM) of the cryogenic LOX, high pressure GOX, and electrolysis approaches across different parameter trades, e.g. mission duration or extravehicular activity frequency, for the single and 10-year mission architectures. For a single nominal mission, high pressure GOX is favored for short missions (< 50 days); cryogenic LOX is favored for a wide-range of mission durations (50 – 270 days); and the electrolysis approach is favored for long missions (> 270 days). However, when considering a 10-year mission architecture, each additional resupply negatively impacts cryogenic LOX due to the additional replacement tankage. Thus, over a 10-year mission, an electrolysis approach, which can provide all life support O2 utilizing solely recovered H2O, appears to be favored over cryogenic LOX. However, a real electrolysis system may need resupplied H2O due to incomplete closure of the air revitalization loop. Thus, the cryogenic LOX approach was compared with the electrolysis approaches utilizing 100% resupplied or 100% recovered H2O to approximate the resupplied to recovered H2O ratio, i.e. the degree of loop closure, where one approach trades over the other. Additionally, gaps were identified, which are expected to affect the viability and trade of cryogenic LOX. These include the development of cryogenic pumps and vaporizers to generate high pressure GOX from LOX as well as understanding payload limitations which can affect O2 resupply.

ECLSS↗
NASA NTRSDate not supplied

Literature Review of Microwave Regeneration of CO2 Capture Sorbent Materials

Utilizing liquid amines for air revitalization applications has been of continued interest for many decades. They have the advantage of being regenerable sorbents with long lifetimes and a high acuteness for carbon dioxide. While their usage on Earth for flue gas and submarines has been explored for some time now, they are still being researched for use in microgravity environments. To succeed in a microgravity environment, such as usage on the International Space Station, an air revitalization system that is reliant on liquid amines as its sorbent must have a low Equivalent System Mass (ESM), among many other performance parameters. One way of optimizing this ESM number is through bringing down the systems power consumption with the use of microwave heating of the liquid sorbent. Microwave heating allows the liquid sorbent to be heated in a much more efficient way when compared to conventional types of heating techniques, such as convection or conduction. This paper will explore the physical mechanisms behind microwave heating, and how it could potentially be applied to a liquid amine carbon capture system, laying out any concerns and future constraints that a system designer may encounter.

Tiago Faria Costa↗
NASA NTRSDate not supplied

Measuring Titanium Heat Pipe Thermal Conductivity for CO2 Adsorption Systems

Current models of the Air-Cooled Temperature Swing Adsorption Compressor (AC-TSAC) being developed for long duration space exploration utilize cooling chambers and heat plates to adsorb and desorb contaminants -13 . To reduce resources required and lower the equivalent system mass (ESM), the study focuses on determining the applications of titanium heat pipe systems to operate both phases, allowing for future research into the development of a heat pipe chamber to hold the zeolite formulations.

multifunctional sorbent↗
NASA NTRSDate not supplied

Development of a Thermal Radiator Optimization Tool with Alternate Coolants

The development of a modeling tool to optimize the design of a thermal radiator based on capacity, rejection temperature, geometry, type and coolant is presented. The radiator size is determined from the desired capacity, environmental conditions and rejection temperature. Flow through the radiator is derived from the overall heat load and prescribed inlet to outlet delta temperature for a given coolant. The tool considers radiator geometry (i.e., tube spacing and diameter, face-sheet thickness, overall size, etc.) and general type (i.e., parallel/manifold versus serpentine) in the optimization. Selection from a handful of potential coolants is also available in the tool which primarily affects necessary tube diameter to maintain turbulent flow and minimize pressure drop. An Equivalent System Mass (ESM) approach is utilized to include pump power in the optimization. Results from the tool and potential future enhancements are also discussed.

Coolant↗
NASA NTRSDate not supplied

Development of a Mechanical Trash Compactor (Mpactor) for Exploration Missions

NASA is developing a mechanical trash compactor (Mpactor), a pneumatically driven trash compaction system, for human space exploration missions to manage trash items such as food packaging, sanitary wipes, clothing, tape, and other types of waste. The compaction system that is under development does not involve heating, water removal, or water recovery which eliminates the need for complex ancillary hardware required to manage gaseous effluents, as well as condense the water vapor. A system called the Crew Exploration Vehicle (CEV) Compactor, intended for use in the micro-gravity environment, was developed at NASA Ames Research Center for the Constellation Program 2007. This CEV Compactor is being further developed and renamed the Mpactor (short for mechanical trash compactor). The Mpactor is being used to gather engineering data that will allow engineers and planners to scale a future design for different mission scenarios and to evaluate different bag materials and designs that help retain compaction.

transit vehicle↗
NASA NTRS2010

Trade Study for 9 kW Water Membrane Evaporator

Sublimators have been proposed and used in spacecraft for heat rejection. Sublimators are desirable heat rejection devices for short duration use because they can transfer large amounts of heat using little mass and are self-regulating devices. Sublimators reject heat into space by freezing water inside a porous substrate, allowing it to sublimate into vapor, and finally venting it into space. The state of the art thermal control system in orbiting spacecraft is a two loop, two fluid system. The external coolant loop typically uses a toxic single phase fluid that acquires heat from the spacecraft and rejects most of it via a radiator. The sublimator functions as a transient topper for orbiting spacecraft during day pass periods when radiator efficiency decreases. The sublimator interfaces with the internal loop through a built in heat exchanger. The internal loop fluid is non-toxic and is typically a propylene glycol and water solution with inhibitors to prevent corrosion with aluminum fins of the heat exchangers. Feedwater is supplied from a separate line to the sublimator to maintain temperature control of the cabin and vehicle hardware. Water membrane evaporators have been developed for spacecraft and spacesuits. They function similar to a sublimator but require a backpressure valve which could be actuated for this application with a simple fully open or fully closed modes. This technology would be applied to orbital thermal control (lunar or planetary). This paper details a trade study showing that evaporators would greatly reduce the consumable that is used, effectively wasted, by sublimators during start up and shut down during the topping phases of each orbit. State of the art for 9 kW sublimators reject about 870 W per kilogram of mass and 1150 W per liter of volume. If water with corrosion inhibitors is used the evaporators would be about 80% of the mass and volume of the equivalent system. The size and mass increases to about 110% if the internal fluid is 50% propylene glycol/50% water. The true benefit comes from the backpressure valve, that prevents the cyclical shutdown/startup loss of the sublimator and amounts to as much as 0.85 kg per orbit.

Bue, Grant C.↗
DOE OSTI2020

High Fidelity Simulations of a Methane-Oxygen Rotating Detonation Rocket Engine

The rotating detonation combustor (RDC) is an important realization of pressure gain combustion for practical applications. The RDC system is characterized by a highly unsteady flow field, with multiple reflected pressure waves following the strong detonations and the existence of partially-burnt gases entrained within the detonation annulus during the injector recovery process. While experimental efforts have provided macroscopic properties of an RDC with a given fuel-oxidizer mixture and geometry, limited optical access and reduced lifespan of flow-field diagnostic equipment constrain understanding of the mechanisms that lend to wave stability, controllability, and sustainability. To this end, high-fidelity numerical simulations of the Air Force Research Laboratory (AFRL) methane-oxygen rotating detonation rocket engine (RDRE) are performed to provide detailed insight into the detonation physics and anomalous behavior within the combustor. The results highlight the merit of high-fidelity numerical studies in analyzing the combustion processes and injector dynamics of a complex combustion system. The simulations capture the general wave behavior observed within the experimental studies. Three different operating conditions, with varying global equivalence ratio and system mass flow rate, are considered: (1) baseline configuration, (2) high equivalence ratio, and (3) high mass flow rate. The high mass flow rate case results in increased detonation wave strength whereas the high equivalence ratio results in increased deflagrative heat release. The high equivalence ratio case results in chaotic behavior with two pairs of counter-rotating detonation waves, whereas a single pair of co-rotating waves is observed in the baseline and high mass flow rate cases. Further, parasitic combustion of the reactant mixture ahead of the detonation wave diminishes its strength and reduces propagation speed to approximately 50-60\% of the Chapman-Jouguet (CJ) condition. In this work, a detailed analysis of the detonation wave behavior, injector dynamics, and the statistics of combustion parameters are presented.

Prakash, Supraj↗
DOE OSTI2021

Dirac systems with magnetic field and position-dependent mass: Darboux transformations and equivalence with generalized Dirac oscillators

Highlights: • We construct Darboux transformations for Dirac systems with position-dependent mass. • Our systems are either Dirac oscillators or coupled to a magnetic field. • We use our Darboux transformation to construct new solvable Dirac systems. We construct a Darboux transformation for a class of two-dimensional Dirac systems at zero energy. Our starting equation features a position-dependent mass, a matrix potential, and an additional degree of freedom that can be interpreted either as a magnetic field perpendicular to the plane or a generalized Dirac oscillator interaction. We obtain a number of Darboux-transformed Dirac equations for which the zero energy solutions are exactly known.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗
DOE OSTI2020

The Most Metal-poor Stars in Omega Centauri (NGC 5139)

The most massive and complex globular clusters in the Galaxy are thought to have originated as the nuclear cores of now tidally disrupted dwarf galaxies, but the connection between globular clusters and dwarf galaxies is tenuous with the M54/Sagittarius system representing the only unambiguous link. The globular cluster Omega Centauri (ω Cen) is more massive and chemically diverse than M 54, and is thought to have been the nuclear star cluster of either the Sequoia or Gaia-Enceladus galaxy. Local Group dwarf galaxies with masses equivalent to these systems often host significant populations of very metal-poor stars ([Fe/H] < −2.5), and one might expect to find such objects in ω Cen. Using high-resolution spectra from Magellan-M2FS, we detected 11 stars in a targeted sample of 395 that have [Fe/H] ranging from −2.30 to −2.52. These are the most metal-poor stars discovered in the cluster, and are five times more metal-poor than ω Cen’s dominant population. However, these stars are not so metal-poor as to be unambiguously linked to a dwarf galaxy origin. The cluster’s metal-poor tail appears to contain two populations near [Fe/H] ∼ −2.1 and −2.4, which are very centrally concentrated but do not exhibit any peculiar kinematic signatures. Several possible origins for these stars are discussed.

79 ASTRONOMY AND ASTROPHYSICS↗
DOE OSTI2022

Super‐Droplet Method to Simulate Lagrangian Microphysics of Nuclear Fallout in a Homogeneous Cloud

Abstract Nuclear detonations produce hazardous local and global particles or fallout. Predicting fallout size, chemical components, and location is necessary to inform officials and determine immediate guidance for the public. However, existing nuclear detonation fallout models prescribe the particle size distributions based on limited observations. In this work, we apply the super‐droplet method, which is a numerical modeling technique developed for cloud microphysics, to simulate size distributions of particles in a mushroom cloud formed post‐detonation of a nuclear device. We model fallout formation and evolution with homogeneous nucleation and condensation of a single species and a Monte Carlo coagulation algorithm. We verify the numerical methods representing coagulation and condensation processes against analytical test problems. Additionally, we explore several scenarios for the integral system mass and yield in equivalent kilotons (kt) of TNT (trinitrotoluene). The fallout size distribution median diameter d pg follows a scaling law based on the integral system mass m v 0 kg and yield Y kt: nm. We test the effect of cloud turbulence, enhanced nucleation and growth, and vapor volatility with a sensitivity study. The range in median diameter predictions for simulations of historical tests performed over the Pacific encompass the measurements of particles sampled from the cloud caps. Predicted median particle size ranges up to 217, 123, 86, and 35 nm for historical tests with yields of 0.2, 0.7, 2, and 10 Mt, respectively. This work can be expanded in many different directions to build a more predictive model for fallout formation.

45 MILITARY TECHNOLOGY, WEAPONRY, AND NATIONAL DEF↗
NASA NTRSDate not supplied

Equivalent Mass Benefits from Employing Vapor Compression Refrigeration on Spacecraft

Thermoelectric cooling, reversed Brayton cycles and Stirling cycle devices are common cooling technologies employed on spacecraft where temperatures below the radiator temperature are needed. Although used less often in microgravity, vapor compression refrigeration was proposed for microgravity applications as early as the 1970s. Researchers proposed the vapor compression cycle over other cooling technologies for its superior energy efficiency in the typical refrigerator/freezer and air-conditioning temperature range. The higher energy efficiency should alleviate the mass penalty of the thermal system, conceivably both for satellites and manned spacecraft. Despite numerous propositions, the literature lacks a discussion of the absolute equivalent mass benefit achievable and the value of it relative to the total mass of the thermal system of the spacecraft or the spacecraft itself. This paper presents the equivalent mass benefits of employing a vapor compression system over other cooling technologies using a mass penalty factor for power consumption. Additionally, a prototype vapor compression cooler sized to fit into an ISS locker is presented.

refrigerator↗
NASA NTRS2018

Design of Distributed Mail-Slot Propulsion System on a Hybrid Wingbody Aircraft

This paper deals with designing a thrust distribution strategy when a Turboelectric Distributed Propulsion (TeDP) system of 16 embedded propulsors is installed on an aerodynamically optimized hybrid wing-body configuration. This HWB previously designed to satisfy conditions of trim, longitudinally static stability and specific cargo space is employed as the baseline configuration for the current study of seeking an optimal propulsion/power system. According to the nature of the entrance flow condition for each distributed propulsion passage in hybrid wing-body aircraft, the ingested boundary layer thickness differs and results in different propulsive reaction. An optimal distribution of thrust and power output is determined by how the system utilizes the propulsive characteristics of each passage. The design space and the number of design variables are selected and described accordingly. An actuator disk model is employed to model thrust generation and shaft power from the propulsor. To carry out the optimization of the propulsion/power system on a computationally expensive CFD model, a Kriging method in conjunction with a Genetic Algorithm (GA) is applied. Throughout the design process, the propulsion performances of the sampled propulsion/power system are analyzed and compared to those of a clean flow engine. The performance metrics includes mass flow rate, fan pressure ratio besides the thrust and shaft power. Minimization of total shaft power from the distributed engine is performed at multiple thrust levels. The benefit of boundary layer ingestion propulsion system is quantified via comparison of thrust equivalent, shaft power and mass flow equivalent clean flow engines with CFD based system design.

Hybrid Wingbody↗