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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 271 records · Page 15
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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 NTRS1974

Mass transfer in a 1370 C (2500 F) lithium thermal convection loop

Experimental results from a test to evaluate interstitial element mass transfer effects on T-111, ASTAR 811C, and ASTAR 1211C after 5000 hours in flowing lithium at 1370 C (2500 F) are presented. No gross corrosion effects were observed. However, hafnium and nitrogen transfer to cooler regions within the loop were noted. Oxygen was in general removed from test specimens, but there was no evidence to indicate that it was a major factor in the mass transfer process. Carbon and hydrogen transfer were not detected.

Scheuermann, C. M.↗
NASA NTRS2008

Investigation of Condensing Ice Heat Exchangers for MTSA Technology Development

Metabolic heat regenerated Temperature Swing Adsorption (MTSA) technology is being developed for thermal, carbon dioxide (CO2) and humidity control for a Portable Life Support Subsystem (PLSS). Metabolically-produced CO2 present in the ventilation gas of a PLSS is collected using a CO2selective adsorbent via temperature swing adsorption. The temperature swing is initiated through cooling to well below metabolic temperatures. Cooling is achieved with a sublimation heat exchanger using water or liquid carbon dioxide (LCO2) expanded below sublimation temperature when exposed to low pressure or vacuum. Subsequent super heated vapor, as well as additional coolant, is used to further cool the astronaut. The temperature swing on the adsorbent is then completed by warming the adsorbent with a separate condensing ice heat exchanger (CIHX) using metabolic heat from moist ventilation gas. The condensed humidity in the ventilation gas is recycled at the habitat. The water condensation from the ventilation gas is a significant heat transfer mechanism for the warming of the adsorbent bed because it represents as much as half of the energy potential in the moist ventilation gas. Designing a heat exchanger to efficiently transfer this energy to the adsorbent bed and allow the collection of the water is a challenge since the CIHX will operate in a temperature range from 210K to 280K. The ventilation gas moisture will first freeze and then thaw, sometimes existing in three phases simultaneously. A NASA Small Business Innovative Research (SBIR) Phase 1 contract was performed to investigate condensing and icing as applied to MTSA to enable higher fidelity modeling and assess the impact of geometry variables on CIHX performance for future CIHX design optimization. Specifically, a design tool was created using analytical relations to explore the complex, interdependent design space of a condensing ice heat exchanger. Numerous variables were identified as having nontrivial contributions to performance such as hydraulic diameter, heat exchanger effectiveness, ventilation gas mass flow rate and surface roughness. Using this tool, four test articles were designed and manufactured to map to a full MTSA subassembly (the adsorbent bed, the sublimation heat exchanger for cooling and the condensing ice heat exchanger for warming). The design mapping considered impacts due to CIHX geometry as well as subassembly impacts such as thermal mass and thermal resistance through the adsorbent bed. The test articles were tested at simulated PLSS ventilation loop temperature, moisture content and subambient pressure. Ice accumulation and melting were observed. Data and test observations were analyzed to identify drivers of the condensing ice heat exchanger performance. This paper will discuss the analytical models, the test article designs, and testing procedures. Testing issues will be discussed to better describe data and share lessons learned. Data analysis and subsequent conclusions will be presented.

Padilla, Sebastian↗
NASA NTRS2014

Mass Dependency of Isotope Fractionation of Gases Under Thermal Gradient and Its Possible Implications for Planetary Atmosphere Escaping Process

Physical processes that unmix elements/isotopes of gas molecules involve phase changes, diffusion (chemical or thermal), effusion and gravitational settling. Some of those play significant roles for the evolution of chemical and isotopic compositions of gases in planetary bodies which lead to better understanding of surface paleoclimatic conditions, e.g. gas bubbles in Antarctic ice, and planetary evolution, e.g. the solar-wind erosion induced gas escaping from exosphere on terrestrial planets.. A mass dependent relationship is always expected for the kinetic isotope fractionations during these simple physical processes, according to the kinetic theory of gases by Chapman, Enskog and others [3-5]. For O-bearing (O16, -O17, -O18) molecules the alpha O-17/ alpha O-18 is expected at 0.5 to 0.515, and for S-bearing (S32,-S33. -S34, -S36) molecules, the alpha S-33/ αlpha S-34 is expected at 0.5 to 0.508, where alpha is the isotope fractionation factor associated with unmixing processes. Thus, one isotope pair is generally proxied to yield all the information for the physical history of the gases. However, we recently] reported the violation of mass law for isotope fractionation among isotope pairs of multiple isotope system during gas diffusion or convection under thermal gradient (Thermal Gradient Induced Non-Mass Dependent effect, TGI-NMD). The mechanism(s) that is responsible to such striking observation remains unanswered. In our past studies, we investigated polyatomic molecules, O2 and SF6, and we suggested that nuclear spin effect could be responsible to the observed NMD effect in a way of changing diffusion coefficients of certain molecules, owing to the fact of negligible delta S-36 anomaly for SF6.. On the other hand, our results also showed that for both diffusion and convection under thermal gradient, this NMD effect is increased by lower gas pressure, bigger temperature gradient and lower average temperature, which indicate that the nuclear spin effect may not be the significant contributor as the energies involved in the hyperfine effect are much smaller than those with molecular collisions, especially under convective conditions.

Sun, Tao↗
NASA NTRS2009

Advanced Fuel Cell System Thermal Management for NASA Exploration Missions

The NASA Glenn Research Center is developing advanced passive thermal management technology to reduce the mass and improve the reliability of space fuel cell systems for the NASA exploration program. An analysis of a state-of-the-art fuel cell cooling systems was done to benchmark the portion of a fuel cell system s mass that is dedicated to thermal management. Additional analysis was done to determine the key performance targets of the advanced passive thermal management technology that would substantially reduce fuel cell system mass.

Burke, Kenneth A.↗
NASA NTRSDate not supplied

Coupling a Computational Fluid Dynamics Model to a Spacecraft Thermal System Model for the DraMS Instrument Thermal Analysis

The Dragonfly Mass Spectrometer (DraMS) is an instrument on the Dragonfly mission, which will spend 7 years in deep space cruise before landing and operating on the surface of Titan. Vacuum thermal analyses are required for deep space cruise, and convection analyses are required for the Titan surface operations. Model exchanges across multiple thermal teams are needed for all phases of the mission. For DraMS, Thermal Desktop® (TD) has been the main thermal analytical tool of choice due to its capability in modeling complex thermal systems with relatively low computational power and for its availability across thermal teams. However, TD does not have computational fluid dynamics (CFD) capability and struggles to accurately capture complex convective behavior. DraMS has fans operating in tandem and gas flow behaviors are not easily predicted due to its complex flow paths. CFD software, such as Fluent, can model and predict such complex flow behaviors, but CFD models are computationally expensive, and its workflow processes are not tailored towards simulating large and complex systems. Therefore, a coupled modeling approach was chosen for DraMS: A TD model was used for simulating all the conductive, radiative, and source terms, while a Fluent CFD model was added on, as needed, to the TD model to provide the convective boundary conditions using the System Coupling software. The coupling software allows the TD and Fluent models to communicate data and arrive at a co-solved and co-converged solution. Furthermore, Thermal Iso-value Exchange (TIE) method was developed to facilitate and improve the TD-Fluent data exchange process. This paper will discuss the analytical studies that were done to verify the accuracy and usability of the coupled approach and the challenges associated, which lead to the development of the TIE approach. DraMS thermal design and co-solved analysis results will also be discussed.

Heat transfer↗
NASA NTRSDate not supplied

Coupling a Computational Fluid Dynamics (CFD) Model to a Spacecraft Thermal System Model for the DraMS Instrument Thermal Analysis

The Dragonfly Mass Spectrometer (DraMS) is an instrument on the Dragonfly mission, which will spend 7 years in deep space cruise before landing and operating on the surface of Titan. Vacuum thermal analyses are required for deep space cruise, and convection analyses are required for the Titan surface operations. Model exchanges across multiple thermal teams are needed for all phases of the mission. For DraMS, Thermal Desktop (TD) has been the main thermal analytical tool of choice due to its capability in modeling complex thermal systems with relatively low computational power and for its availability across thermal teams. However, TD does not have computational fluid dynamics (CFD) capability and struggles to accurately capture complex convective behavior. DraMS has fans operating in tandem and gas flow behaviors are not easily predicted due to its complex flow paths. CFD software, such as Fluent, can model and predict such complex flow behaviors, but CFD models are computationally expensive, and its workflow processes are not tailored towards simulating large and complex systems. Therefore, a coupled modeling approach was chosen for DraMS: A TD model was used for simulating all the conductive, radiative, and source terms, while a Fluent CFD model was added on, as needed, to the TD model to provide the convective boundary conditions using the System Coupling software. The coupling software allows the TD and Fluent models to communicate data and arrive at a co-solved and co-converged solution. Furthermore, Thermal Iso-value Exchange (TIE) method was developed to facilitate and improve the TD-Fluent data exchange process. This paper will discuss the analytical studies that were done to verify the accuracy and usability of the coupled approach and the challenges associated, which lead to the development of the TIE approach. DraMS thermal design and co-solved analysis results will also be discussed.

heat transfer↗
NASA NTRS2008

Phoenix Lander's Thermal Evolved Gas Analyzer: Differential Scanning Calorimeter and Mass Spectrometer Database Development

The Mars Scout Phoenix lander will land in the north polar region of Mars in May, 2008. One objective of the Phoenix lander is to search for evidence of past life in the form of molecular organics that may be preserved in the subsurface soil. The Thermal Evolved Gas Analyzer (TEGA) was developed to detect these organics by coupling a simultaneous differential thermal analyzer (SDTA) with a mass spectrometer. Martian soil will be heated to approx.1000 C and potential organic decomposition products such as CO2, CH4 etc. will be examined for with the MS. TEGA s SDTA will also assess the presence of endothermic and exothermic reactions that are characteristic of soil organics and minerals as the soil is heated. The MS in addition to detecting organic decompositon products, will also assess the levels of soil inorganic volatiles such as H2O, SO2, and CO2. Organic detection has a high priority for this mission; however, TEGA has the ability to provide valuable insight into the mineralogical composition of the soil. The overall goal of this work is to develop a TEGA database of minerals that will serve as a reference for the interpretation of Phoenix-TEGA. Previous databases for the ill-fated Mars Polar Lander (MPL)-TEGA instrument only went to 725 C. Furthermore, the MPL-TEGA could only detect CO2 and H2O while the Phoenix-TEGA MS can examine up to 144 atomic mass units. The higher temperature Phoenix-TEGA SDTA coupled with the more capable MS indicates that a higher temperature database is required for TEGA interpretation. The overall goal of this work is to develop a differential scanning calorimeter (DSC) database of minerals along with corresponding MS data of evolved gases that can used to interpret TEGA data during and after mission operations. While SDTA and DSC measurement techniques are slightly different (SDTA does not use a reference pan), the results are fundamentally similar and thus DSC is a useful technique in providing comparative data for the TEGA database. The objectives of this work is to conduct DSC and MS analysis up to 1000 C of select minerals that may be found in the martian soil.

Sutter, B.↗
NASA NTRS2020

Massive Galaxies Impede Massive Outflows

A set of 66 3D hydrodynamical simulations explores how galactic stellar mass affects three-phase, starburst-driven outflows. Simulated velocities are compared to two basic analytic models: with and without a gravitational potential. For stellar mass <1010 M ⊙, simulated velocities match those of both analytical models and are unaffected by the potential; above they reduce significantly as expected from the analytic model with gravity. Gravity also affects total outflow mass and each of the three phases differently. Outflow masses in the hot, warm, and cold phases each scale with stellar mass as $\mathrm{log}{M}_{* }\,=$ −0.25, −0.97, and −1.70, respectively. Thus, the commonly used Chevalier & Clegg analytic model should be modified to include gravity when applied to higher-mass galaxies. In particular, using M82 as the canonical galaxy to interpret hydrodynamical simulations of starburst-driven outflows from higher-mass galaxies will underestimate the retarding effect of gravity. Using the analytic model of Johnson & Axford with realistic thermalization efficiency and mass loading, I find that only galaxy masses lesssim10(exp 11.5) M(⊙) can outflow.

Ryan Tanner↗
NASA NTRS1990

X-ray photoelectron and mass spectroscopic study of electron irradiation and thermal stability of polytetrafluoroethylene

Polytetrafluoroethylene (PTFE) was subjected to 3 keV electron bombardment and then heated in vacuum to 300 C. The behavior of the material as a function of radiation dose and temperature was studied by X-ray photoelectron spectroscopy (XPS) of the surface and mass spectroscopy of the species evolved. A quantitative comparison of the radiation dose rate with that in other reported studies showed that, for a given total dose, the damage observed by XPS is greater for higher dose rates. Lightly damaged material heated to 300 C evolved saturated fluorocarbon species, whereas unsaturated fluorocarbon species evolved from heavily damaged material. After heating the heavily damaged material, those features in the XPS that were associated with damage diminished, giving the appearance that the radiation damage annealed. The apparent annealing of the radiation damage was found to be due to the covering of the network by saturated fragments that easily diffused through the decomposed material to the surface region upon heating.

Wheeler, Donald R.↗
NASA NTRS2003

Direct detection of the Yarkovsky effect by radar ranging to ateroid 6489 Golevka

Radar ranging from Arecibo, Puerto Rico, to the 0.5-kilometer near-Earth asteroid 6489 Golevka unambiguously reveals a small nongravitational acceleration caused by the anisotropic thermal emission of absorbed sunlight. The magnitude of this perturbation, known as the Yarkovsky effect, is a function of the asteroid's mass and surface thermal characteristics. Direct detection of the Yarkovsky effect on asteroids will help constrain their physical properties, such as bulk density, and refine their orbital paths.

asteriod mass radar orbital paths↗
NASA NTRS2018

A Common Probe Design for Multiple Planetary Destinations

Atmospheric probes have been successfully flown to planets and moons in the solar system to conduct in situ measurements. They include the Pioneer Venus multi-probes, the Galileo Jupiter probe, and Huygens probe. Probe mission concepts to five destinations, including Venus, Jupiter, Saturn, Uranus, and Neptune, have all utilized similar-shaped aeroshells and concept of operations, namely a 45-degree sphere cone shape with high density heatshield material and parachute system for extracting the descent vehicle from the aeroshell. Each concept designed its probe to meet specific mission requirements and to optimize mass, volume, and cost. At the 2017 International Planetary Probe Workshop (IPPW), NASA Headquarters postulated that a common aeroshell design could be used successfully for multiple destinations and missions. This "common probe" design could even be assembled with multiple copies, properly stored, and made available for future NASA missions, potentially realizing savings in cost and schedule and reducing the risk of losing technologies and skills difficult to sustain over decades. Thus the NASA Planetary Science Division funded a study to investigate whether a common probe design could meet most, if not all, mission needs to the five planetary destinations with extreme entry environments. The Common Probe study involved four NASA Centers and addressed these issues, including constraints and inefficiencies that occur in specifying a common design. Study methodology: First, a notional payload of instruments for each destination was defined based on priority measurements from the Planetary Science Decadal Survey. Steep and shallow entry flight path angles (EFPA) were defined for each planet based on qualification and operational g-load limits for current, state-of-the-art instruments. Interplanetary trajectories were then identified for a bounding range of EFPA. Next, 3-degrees-of-freedom simulations for entry trajectories were run using the entry state vectors from the interplanetary trajectories. Aeroheating correlations were used to generate stagnation point convective and radiative heat flux profiles for several aeroshell shapes and entry masses. High fidelity thermal response models for various Thermal Protection System (TPS) materials were used to size stagnation-point thicknesses, with margins based on previous studies. Backshell TPS masses were assumed based on scaled heat fluxes from the heatshield and also from previous mission concepts. Presentation: We will present an overview of the study scope, highlights of the trade studies and design driver analyses, and the final recommendations of a common probe design and assembly. We will also indicate limitations that the common probe design may have for the different destinations. Finally, recommended qualification approaches for missions will be presented.

Hwang, H. H.↗
NASA NTRS2018

A Common Probe Design for Multiple Planetary Destinations

Atmospheric probes have been successfully flown to planets and moons in the solar system to conduct in-situ measurements. They include the Pioneer Venus multi-probes, the Galileo Jupiter probe, and Huygens probe. Probe mission concepts to five destinations, including Venus, Jupiter, Saturn, Uranus, and Neptune, have all utilized similar-shaped aeroshells and concept of operations, namely a 45 deg sphere cone shape with high density heatshield material and parachute system for extracting the descent vehicle from the aeroshell. Each concept designed its probe to meet specific mission requirements and to optimize mass, volume, and cost. At the 2017 IPPW, NASA Headquarters postulated that a common aero-shell design could be used successfully for multiple destinations and missions. This "common probe" design could even be assembled with multiple copies, properly stored, and made available for future NASA missions, potentially realizing savings in cost and schedule and reducing the risk of losing technologies and skills difficult to sustain over decades. Thus the NASA Planetary Science Division funded a study to investigate whether a common probe design could meet most, if not all, mission needs to the five planetary destinations with extreme entry environments. The Common Probe study involved four NASA Centers and addressed these issues, including constraints and inefficiencies that occur in specifying a common design.Study methodology: First, a notional payload of instruments for each destination was defined based on priority measurements from the Planetary Science Decadal Survey. Steep and shallow entry flight path angles (EFPA) were defined for each planet based on qualification and operational g-load limits for current, state-of-the-art instruments. Interplanetary trajectories were then identified for a bounding range of EFPA. Next, 3-DoF simulations for entry trajectories were run using the entry state vectors from the interplanetary trajectories. Aeroheating correlations were used to generate stagnation point convective and radiative heat flux profiles for several aeroshell shapes and entry masses. High fidelity thermal response models for various TPS materials were used to size stagnation point thicknesses, with margins based on previous studies. Backshell TPS masses were assumed based on scaled heat fluxes from the heatshield and also from previous mission concepts.

Hwang, Helen H.↗
NASA NTRS2026

Simulating the Lunar Thermal Environment for the Surface-Deployed LEMS Artemis III TVAC Test

The Lunar Environment Monitoring Station (LEMS) is an autonomous, survive-the lunar-night seismic suite to be deployed on the Lunar surface by the Artemis III crew and designed to operate continuously for two years. It will see the extremes of the Lunar south pole thermal environment where surface temperatures range from -200°C to +20°C and where nighttime duration is at least 354 hours. Given power and mass constraints, the thermal system is limited to 2.5 Watts of heat during the lunar night. Presented is the system thermal vacuum test plan and configuration to validate the thermal control system performance in a flight-like environment. The test will measure the nighttime heat leaks for the bus and 18 seismometers to ensure lunar night survival and operability.

Thermal↗
NASA NTRSDate not supplied

Harness Thermal Heat Loss Measurement for a Lunar Surface-Deployed LEMS Artemis III Payload

The Lunar Environment Monitoring Station (LEMS) is an autonomous, survive-the-lunar-night seismic suite to be deployed on the Lunar surface by the Artemis III crew and designed to operate continuously for two years. It will see the extremes of the Lunar south pole thermal environment where surface temperatures range from -200°C to +20°C and where nighttime duration is at least 354 hours. Given power and mass constraints, the thermal system is limited to 2.5 Watts of heat during the lunar night. The electrical harnessing named the Signal and Power Passthrough (SAPP) was designed to minimize heat loss while meeting power and signal integrity requirements. To mitigate risk due to uncertainty associated with the harnessing materials, routing, and tie-downs, a flight-like thermal conductance test was performed to measure the heat loss. The test methodology, results, and model correlation are presented.

TVAC↗
NASA NTRS2026

Simulating the Lunar Thermal Environment for the Surface-Deployed LEMS Artemis III TVAC Test

The Lunar Environment Monitoring Station (LEMS) is an autonomous, survive-the lunar-night seismic suite to be deployed on the Lunar surface by the Artemis III crew and designed to operate continuously for two years. It will see the extremes of the Lunar south pole thermal environment where surface temperatures range from -200°C to +20°C and where nighttime duration is at least 354 hours. Given power and mass constraints, the thermal system is limited to 2.5 Watts of heat during the lunar night. Presented is the system thermal vacuum test plan and configuration to validate the thermal control system performance in a flight-like environment. The test will measure the nighttime heat leaks for the bus and 18 seismometers to ensure lunar night survival and operability.

Thermal↗
NASA NTRSDate not supplied

Harness Thermal Heat Loss Measurement for A Lunar Surface-Deployed LEMS Artemis III Payload

The Lunar Environment Monitoring Station (LEMS) is an autonomous, survive-the-lunar-night seismic suite to be deployed on the Lunar surface by the Artemis III crew and designed to operate continuously for two years. It will see the extremes of the Lunar south pole thermal environment where surface temperatures range from -200°C to +20°C and where nighttime duration is at least 354 hours. Given power and mass constraints, the thermal system is limited to 2.5 Watts of heat during the lunar night. The electrical harnessing named the Signal and Power Passthrough (SAPP) was designed to minimize heat loss while meeting power and signal integrity requirements. To mitigate risk due to uncertainty associated with the harnessing materials, routing, and tie-downs, a flight-like thermal conductance test was performed to measure the heat loss. The test methodology, results, and model correlation are presented.

Thermal↗
NASA NTRSDate not supplied

Harness Thermal Heat Loss Measurement for A Lunar Surface-Deployed LEMS Artemis III Payload

The Lunar Environment Monitoring Station (LEMS) is an autonomous, survive-the-lunar-night seismic suite to be deployed on the Lunar surface by the Artemis III crew and designed to operate continuously for two years. It will see the extremes of the Lunar south pole thermal environment where surface temperatures range from -200°C to +20°C and where nighttime duration is at least 354 hours. Given power and mass constraints, the thermal system is limited to 2.5 Watts of heat during the lunar night. The electrical harnessing named the Signal and Power Passthrough (SAPP) was designed to minimize heat loss while meeting power and signal integrity requirements. To mitigate risk due to uncertainty associated with the harnessing materials, routing, and tie-downs, a flight-like thermal conductance test was performed to measure the heat loss. The test methodology, results, and model correlation are presented.

TVAC↗