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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 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

Variable Conductance Heat Pipes for Radioisotope Stirling Systems

In a Stirling radioisotope system, heat must continually be removed from the GPHS modules, to maintain the GPHS modules and surrounding insulation at acceptable temperatures. Normally, the Stirling convertor provides this cooling. If the Stirling engine stops in the current system, the insulation is designed to spoil, preventing damage to the GPHS, but also ending the mission. An alkali-metal Variable Conductance Heat Pipe (VCHP) was designed to allow multiple stops and restarts of the Stirling engine. A VCHP turns on with a delta T of 30 C, which is high enough to not risk standard ASRG operation but low enough to save most heater head life. This VCHP has a low mass, and low thermal losses for normal operation. In addition to the design, a proof-of-concept NaK VCHP was fabricated and tested. While NaK is normally not used in heat pipes, it has an advantage in that it is liquid at the reservoir operating temperature, while Na or K alone would freeze. The VCHP had two condensers, one simulating the heater head, and the other simulating the radiator. The experiments successfully demonstrated operation with the simulated heater head condenser off and on, while allowing the reservoir temperature to vary over 40 to 120 C, the maximum range expected. In agreement with previous NaK heat pipe tests, the evaporator delta T was roughly 70 C, due to distillation of the NaK in the evaporator.

Anderson, William G.↗
NASA NTRS2011

Radiolytic Cryovolcanism Revisited

Location of Enceladus within the inner magnetosphere of Saturn makes it likely that energetic electron irradiation, ion-neutral chemistry of the emergent cryovolcanic plume gas, and resultant radiolytic modification of surface ice composition could impact long-term evolution of molecular composition in the global ice crust. Thermally-driven convection of ice in the south polar terrain would bring radiolytic oxidant products into contact with subsurface reservoirs of primordial organics on million-year time scales. The chemical reactions leading to CO2 gas production as a primary gas-piston driver of cryovolcanic activity would be exponentially elevated in the warm-ice margins of a heated fluid. The fluid temperature is typically assumed to be below 273 K but could be much higher in a gas-pressurized deep subsurface environment. The Perrier Ocean model has demonstrated how a CO2-loaded fluid could account for the observed jets, while the heat content of the fluid arising from the moon's deeper interior could support high levels of chemical reactivity in the thermal margins. Since mass loss and tidal dissipation arguments do not support continuous activity over billions of years, the activity is likely very episodic so that even low-level energy sources including irradiation-driven radiolysis of surface ices could substantially contribute to the chemical dynamics of the activity apparently now in high phase. A multi-phase thermochemical model, and supporting laboratory measurements of temperature-dependent reaction rates, are needed to investigate these potentially complex processes.

Cooper, John F.↗
NASA NTRS2010

A Study of Fluid Interface Configurations in Exploration Vehicle Propellant Tanks

The equilibrium shape and location of fluid interfaces in spacecraft propellant tanks while in low-gravity is of interest to system designers, but can be challenging to predict. The propellant position can affect many aspects of the spacecraft such as the spacecraft center of mass, response to thruster firing due to sloshing, liquid acquisition, propellant mass gauging, and thermal control systems. We use Surface Evolver, a fluid interface energy minimizing algorithm, to investigate theoretical equilibrium liquid-vapor interfaces for spacecraft propellant tanks similar to those that have been considered for NASA's new class of Exploration vehicles. The choice of tank design parameters we consider are derived from the NASA Exploration Systems Architecture Study report. The local acceleration vector employed in the computations is determined by estimating low-Earth orbit (LEO) atmospheric drag effects and centrifugal forces due to a fixed spacecraft orientation with respect to the Earth or Moon, and rotisserie-type spacecraft rotation. Propellant/vapor interface positions are computed for the Earth Departure Stage and Altair lunar lander descent and ascent stage tanks for propellant loads applicable to LEO and low-lunar orbit. In some of the cases investigated the vapor ullage bubble is located at the drain end of the tank, where propellant management device hardware is often located.

Zimmerli, Gregory A.↗
NASA NTRS2011

Design and Modeling of a Variable Heat Rejection Radiator

Variable Heat Rejection Radiator technology needed for future NASA human rated & robotic missions Primary objective is to enable a single loop architecture for human-rated missions (1) Radiators are typically sized for maximum heat load in the warmest continuous environment resulting in a large panel area (2) Large radiator area results in fluid being susceptible to freezing at low load in cold environment and typically results in a two-loop system (3) Dual loop architecture is approximately 18% heavier than single loop architecture (based on Orion thermal control system mass) (4) Single loop architecture requires adaptability to varying environments and heat loads

fluid flow↗
NASA NTRS2012

Experimental Studies of Carbon Nanotube Materials for Space Radiators

Game ]changing propulsion systems are often enabled by novel designs using advanced materials. Radiator performance dictates power output for nuclear electric propulsion (NEP) systems. Carbon nanotubes (CNT) and carbon fiber materials have the potential to offer significant improvements in thermal conductivity and mass properties. A test apparatus was developed to test advanced radiator designs. This test apparatus uses a resistance heater inside a graphite tube. Metallic tubes can be slipped over the graphite tube to simulate a heat pipe. Several sub ]scale test articles were fabricated using CNT cloth and pitch ]based carbon fibers, which were bonded to a metallic tube using an active braze material. The test articles were heated up to 600 C and an infrared (IR) camera captured the results. The test apparatus and experimental results are presented here.

SanSoucie, MIchael P.↗
NASA NTRS2015

The Transition Region Response to a Coronal Nanoflare: Forward Modeling and Observations in SDO/AIA

The corona and transition region (TR) are fundamentally coupled through the processes of thermal conduction and mass exchange. It is not possible to understand one without the other. Yet the temperature-dependent emissions from the two locations behave quite differently in the aftermath of an impulsive heating event such as a coronal nanoflare. Whereas the corona cools sequentially, emitting first at higher temperatures and then at lower temperatures, the TR is multithermal and the emission at all temperatures responds in unison. We have previously applied the automated time lag technique of Viall & Klimchuk to disk observations of an active region (AR) made by the Atmospheric Imaging Assembly (AIA) on the Solar Dynamics Observatory. Lines of sight passing through coronal plasma show clear evidence for post-nanoflare cooling, while lines of sight intersecting the TR footpoints of coronal strands show zero time lag. In this paper, we use the EBTEL hydrodynamics code to demonstrate that this is precisely the expected behavior when the corona is heated by nanoflares. We also apply the time lag technique for the first time to off-limb observations of an AR. Since TR emission is not present above the limb, the occurrence of zero time lags is greatly diminished, supporting the conclusion that zero time lags measured on the disk are due to TR plasma. Lastly, we show that the "coronal" channels in AIA can be dominated by bright TR emission. When defined in a physically meaningful way, the TR reaches a temperature of roughly 60% the peak temperature in a flux tube. The TR resulting from impulsive heating can extend to 3 MK and higher, well within the range of the "coronal" AIA channels.

Sun: corona↗
NASA NTRS2011

Technique for Configuring an Actively Cooled Thermal Shield in a Flight System

Broad area cooling shields are a mass-efficient alternative to conductively cooled thermal radiation shielding. The shield would actively intercept a large portion of incident thermal radiation and transport the heat away using cryogenic helium gas. The design concept consists of a conductive and conformable surface that maximizes heat transfer and formability. Broad Area Cooled (BAC) shields could potentially provide considerable mass savings for spaceflight applications by eliminating the need for a rigid thermal radiation shield for cryogen tanks. The BAC consists of a network of capillary tubes that are thermally connected to a conductive shield material. Chilled helium gas is circulated through the network and transports unwanted heat away from the cryogen tanks. The cryogenic helium gas is pumped and chilled simultaneously using a specialized pulse-tube cryocooler, which further improves the mass efficiency of the system. By reducing the thermal environment temperature from 300 to 100 K, the radiative heat load on a cryogen tank could be reduced by an order of magnitude. For a cryogenic liquid propellant scenario of oxygen and hydrogen, the boiloff of hydrogen would be significantly reduced and completely eliminated for oxygen. A major challenge in implementing this technology on large tanks is that the BAC system must be easily scalable from lab demonstrations to full-scale missions. Also, the BAC shield must be conformable to complex shapes like spheres without losing the ability to maintain constant temperature throughout. The initial design maximizes thermal conductivity between the capillary tube and the conductive radiation shielding by using thin, corrugated aluminum foil with the tube running transverse to the folds. This configuration has the added benefit of enabling the foil to stretch and contract longitudinally. This allows the BAC to conform to the complex curvature of a cryogen tank, which is key to its success. To demonstrate a BAC shield system with minimal impact to current cryogen tank designs, the shielding must be applied after the final assembly of the tank and supporting structure. One method is to pre-fabricate the shield in long strips. A spool of corrugated aluminum foil with a thermally sunk aluminum capillary running through the center could then be simply wound around the cryogen tanks and encapsulated within the multi-layer insulation (MLI) blanket. Then, on orbit, the BAC would intercept thermal radiation coming in through the MLI and transport it away from the cryogen tanks. An optimization of the design could be done to take into account mass savings from thinner MLI blankets, eliminating solid thermal shields, and ultimately, a reduction in the required cryogen tank size.

Barkfknecht, Peter↗
NASA NTRS2009

Lunar Thermal Wadis and Exploration Rovers: Outpost Productivity and Participatory Exploration

The presentation introduces the concept of a thermal wadi, an engineered source of thermal energy that can be created using native material on the moon or elsewhere to store solar energy for use by various lunar surface assets to survive the extremely cold environment of the lunar night. A principal benefit of this approach to energy storage is the low mass requirement for transportation from Earth derived from the use of the lunar soil, or regolith, as the energy storage medium. The presentation includes a summary of the results of a feasibility study involving the numerical modeling of the performance of a thermal wadi including a manufactured thermal mass, a solar energy reflector, a nighttime thermal energy reflector and a lunar surface rover. The feasibility study shows that sufficient thermal energy can be stored using unconcentrated solar flux to keep a lunar surface rover sufficiently warm throughout a 354 hour lunar night at the lunar equator, and that similar approaches can be used to sustain surface assets during shorter dark periods that occur at the lunar poles. The presentation includes descriptions of a compact lunar rover concept that could be used to manufacture a thermal wadi and could alternatively be used to conduct a variety of high-value tasks on the lunar surface. Such rovers can be produced more easily because the capability for surviving the lunar night is offloaded to the thermal wadi infrastructure. The presentation also includes several concepts for operational scenarios that could be implemented on the moon using the thermal wadi and compact rover concepts in which multiple affordable rovers, operated by multiple terrestrial organizations, can conduct resource prospecting and human exploration site preparation tasks.

Sacksteder, Kurt↗
NASA NTRS1996

Support of gas flowmeter upgrade

A project history review, literature review, and vendor search were conducted to identify a flowmeter that would improve the accuracy of gaseous flow measurements in the White Sands Test Facility (WSTF) Calibration Laboratory and the Hydrogen High Flow Facility. Both facilities currently use sonic flow nozzles to measure flowrates. The flow nozzle pressure drops combined with corresponding pressure and temperature measurements have been estimated to produce uncertainties in flowrate measurements of 2 to 5 percent. This study investigated the state of flowmeter technology to make recommendations that would reduce those uncertainties. Most flowmeters measure velocity and volume, therefore mass flow measurement must be calculated based on additional pressures and temperature measurement which contribute to the error. The two exceptions are thermal dispersion meters and Coriolis mass flowmeters. The thermal dispersion meters are accurate to 1 to 5 percent. The Coriolis meters are significantly more accurate, at least for liquids. For gases, there is evidence they may be accurate to within 0.5 percent or better of the flowrate, but there may be limitations due to inappropriate velocity, pressure, Mach number and vibration disturbances. In this report, a comparison of flowmeters is presented. Candidate Coriolis meters and a methodology to qualify the meter with tests both at WSTF and Southwest Research Institute are recommended and outlined.

Waugaman, Dennis↗
NASA NTRS1972

Laser mass spectrometer

Development of time of flight mass spectrometer using laser powered vaporization source is discussed. Operation of equipment to measure thermal velocities of individual mass species is described. Illustration of test equipment installation and detailed line drawing of ion producing equipment are provided.

Lincoln, K. A.↗
NASA NTRS2011

X-Ray Spectra from MHD Simulations of Accreting Black Holes

We present new global calculations of X-ray spectra from fully relativistic magneto-hydrodynamic (MHO) simulations of black hole (BH) accretion disks. With a self consistent radiative transfer code including Compton scattering and returning radiation, we can reproduce the predominant spectral features seen in decades of X-ray observations of stellar-mass BHs: a broad thermal peak around 1 keV, power-law continuum up to >100 keV, and a relativistically broadened iron fluorescent line. By varying the mass accretion rate, different spectral states naturally emerge: thermal-dominant, steep power-law, and low/hard. In addition to the spectral features, we briefly discuss applications to X-ray timing and polarization.

Schnittman, Jeremy D.↗
NASA NTRS1987

What is the essential physics of mass loss from late-type stars?

A review is given of the data concerning mass loss from late-type stars. The major classes of mass-loss mechanisms (thermally-driven winds, radiatively-driven winds, and wave-driven winds) are discussed, and the empirical mass-loss rates and other data are compared with theoretical predictions for any of these mechanisms acting alone. It is likely that several mechanisms act together to produce the large mass-loss rates in the Mira and nonpulsating M supergiants. Studies of the solar atmosphere suggest that thermal bifurcation driven by molecular condensation instabilities may play a critical role in cooling the atmospheres of luminous cool stars and forming siicate dust. It is possible that several metastable modes of atmospheric structure may exist for a given set of stellar parameters.

Linsky, Jeffrey L.↗
NASA NTRS1997

Adsorption and solvation of ethanol at the water liquid-vapor interface: a molecular dynamics study

The free energy profiles of methanol and ethanol at the water liquid-vapor interface at 310K were calculated using molecular dynamics computer simulations. Both alcohols exhibit a pronounced free energy minimum at the interface and, therefore, have positive adsorption at this interface. The surface excess was computed from the Gibbs adsorption isotherm and was found to be in good agreement with experimental results. Neither compound exhibits a free energy barrier between the bulk and the surface adsorbed state. Scattering calculations of ethanol molecules from a gas phase thermal distribution indicate that the mass accommodation coefficient is 0.98, and the molecules become thermalized within 10 ps of striking the interface. It was determined that the formation of the solvation structure around the ethanol molecule at the interface is not the rate-determining step in its uptake into water droplets. The motion of an ethanol molecule in a water lamella was followed for 30 ns. The time evolution of the probability distribution of finding an ethanol molecule that was initially located at the interface is very well described by the diffusion equation on the free energy surface.

NASA Discipline Exobiology↗
NASA NTRS1980

Mass reduction for advanced winged entry vehicles through integrated thermostructural-trajectory design

This study examines the impact of aerothermo-structure design integration on thermal protection system (TPS) mass for advanced winged entry vehicles. Four basic categories of TPS are considered: external insulation, metallic hot structures, metallic standoff, and hybrid systems. A variety of concepts that fit in these basic categories are examined. Entry trajectories tailored specifically to the characteristics of each TPS concept are generated. An aerodynamic heating program is used with the appropriately tailored trajectories to determine the centerline TPS requirements for each concept. Included in the investigation is the impact on the combined TPS-structure mass of structural materials which maintain their integrity to higher temperatures. Results indicate that significant mass reduction can be realized through appropriate aerothermo-structure design integration. An assessment is made of the relative merits of the TPS/structure/trajectory combinations considered in the study.

Wurster, K. E.↗
NASA NTRS2016

Optimization of NTP System Truss to Reduce Radiation Shield Mass

The benefits of nuclear thermal propulsion are numerous and relevant to the current NASA mission goals involving but not limited to the crewed missions to mars and the moon. They do however also present new and unique challenges to the design and logistics of launching/operating spacecraft. One of these challenges, relevant to this discussion, is the significant mass of the shielding which is required to ensure an acceptable radiation environment for the spacecraft and crew. Efforts to reduce shielding mass are difficult to accomplish from material and geometric design points of the shield itself, however by increasing the distance between the nuclear engines and the main body of the spacecraft the required mass of the shielding is lessened considerably. The mass can be reduced significantly per unit length, though any additional mass added by the structure to create this distance serves to offset those savings, thus the design of a lightweight structure is ideal. The challenges of designing the truss are bounded by several limiting factors including; the loading conditions, the capabilities of the launch vehicle, and achieving the ideal truss length when factoring for the overall mass reduced. Determining the overall set of mass values for a truss of varying length is difficult since to maintain an optimally designed truss the geometry of the truss or its members must change. Thus the relation between truss mass and length for these loading scenarios is not linear, and instead has relation determined by the truss design. In order to establish a mass versus length trend for various truss designs to compare with the mass saved from the shield versus length, optimization software was used to find optimal geometric properties that still met the design requirements at established lengths. By solving for optimal designs at various lengths, mass trends could be determined. The initial design findings show a clear benefit to extending the engines as far from the main structure of the spacecraft as the launch vehicle's payload volume would allow when comparing mass savings verse the additional structure.

Scharber, Luke L.↗
NASA NTRS1983

Radiative equilibrium model of Titan's atmosphere

The present global radiative equilibrium model for the Saturn satellite Titan is restricted to the two-stream approximation, is vertically homogeneous in its scattering properties, and is spectrally divided into one thermal and two solar channels. Between 13 and 33% of the total incident solar radiation is absorbed at the planetary surface, and the 30-60 ratio of violet to thermal IR absorption cross sections in the stratosphere leads to the large temperature inversion observed there. The spectrally integrated mass absorption coefficient at thermal wavelengths is approximately constant throughout the stratosphere, and approximately linear with pressure in the troposphere, implying the presence of a uniformly mixed aerosol in the stratosphere. There also appear to be two regions of enhanced opacity near 30 and 500 mbar.

Samuelson, R. E.↗