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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 829 records · Page 46
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Effect of HIPing on the effective thermal conductivity/diffusivity and the interfacial thermal conductance of uniaxial SiC fibre-reinforced RBSN

Hot isostatic pressing (HIPing) was found to increase the thermal diffusivity/conductivity of uniaxial silicon carbide fiber-reinforced reaction-bonded silicon nitride (RBSN) matrix composites, as the result of the densification of the matrix, the increase in the grain size of the silicon carbide and the improved thermal contact between the fibers and the matrix. Transverse to the fiber direction the thermal diffusivity/conductivity was found to be a function of the surrounding gaseous atmosphere due to the access of the gas phase to the fiber-matrix interface, which was facilitated by the existence of an interfacial gap due to the thermal expansion mismatch between the fibers and the matrix. The interfacial conductance was found to exhibit a strong positive temperature dependence as the result of the closure of the interfacial gap with increasing temperature.

Bhatt, H.↗
NASA NTRS2011

End-of-Life Optical Property Predictions of White Conductive Thermal Control Coatings through Analysis of On-Orbit and Ground Based Testing Data

New system requirements pertaining to thermal optical properties and coating electrical properties are commonly specified on non-low earth orbit missions. An increasing number of projects are specifying coatings with a surface resistivity of less than lE-9 ohm/square to mitigate electrostatic charge buildup events over a range of operational temperatures. There are a limited number of coatings that. meet these electrical property requirements while having flight derived optical properties in representative environments. Goddard Space Flight Center Code 546, Contamination and Thermal Coatings Group has recently explored the variety of electrically conductive white coatings available through domestic vendors to evaluate properties to meet project requirements in a geostationary orbit. The lack of significant flight data in representative environments required the careful selection of samples in ground based tests to establish end of life thermal properties. Attention must be given to the origin and pedigree of samples used on past on-orbit experiments to insure that the present formulations for the materials are similar and will react in similar manner.

Hasegawa, Mark↗
NASA NTRS1998

Measuring Thermal Conductivity and Moisture Absorption of Cryo-Insulation Materials

NASA is seeking to develop thermal insulation material systems suitable for withstanding both extremely high temperatures encountered during atmospheric re-entry heating and aero- braking maneuvers, as well as extremely low temperatures existing in liquid fuel storage tanks. Currently, materials used for the high temperature insulation or Thermal Protection System (TPS) are different from the low temperature, or cryogenic insulation. Dual purpose materials are necessary to the development of reusable launch vehicles (RLV). The present Space Shuttle (or Space Transportation System, STS) employs TPS materials on the orbiter and cryo-insulation materials on the large fuel tank slung under the orbiter. The expensive fuel tank is jettisoned just before orbit is achieved and it burns up while re-entering over the Indian Ocean. A truly completely reusable launch vehicle must store aR cryogenic fuel internally. The fuel tanks will be located close to the outer surface. In fact the outer skin of the craft will probably also serve as the fuel tank enclosure, as in jet airliners. During a normal launch the combined TPS/cryo-insulation system will serve only as a low temperature insulator, since aerodynamic heating is relatively minimal during ascent to orbit. During re-entry, the combined TPS/cryo-insulation system will serve only as a high temperature insulator, since all the cryogenic fuel will have been expended in orbit. However, in the event of an.aborted launch or a forced/emergency early re-entry, the tanks will still contain fuel, and the TPS/cryo-insulation will have to serve as both low and high temperature insulation. Also, on long duration missions, such as to Mars, very effective cryo-insulation materials are needed to reduce bod off of liquid propellants, thereby reducing necessary tankage volume, weight, and cost. The conventional approach to obtaining both low and high temperature insulation, such as is employed for the X-33 and X-34 spacecraft, is to use separate TPS and cryo-insulation materials, which are connected by means of adhesives or stand-offs (spacers). Three concepts are being considered: (1) the TPS is bonded directly to the cryo-insulation which, in turn, is bonded to the exterior of the tank, (2) stand-offs are used to make a gap between the TPS and the cryo-insulation, which is bonded externally to the tank, (3) TPS is applied directly or with stand-offs to the exterior so the tank, and cryo-insulation is applied directly to the interior of the tank. Many potential problems are inherent in these approaches. For example, mismatch between coefficients of thermal expansion of the TPS and cryo-insulation, as well as aerodynamic loads, could lead to failure of the bond. Internal cryo-insulation must be prevent from entering the sump of the fuel turbo-pump. The mechanical integrity of the stand-off structure (if used) must withstand multiple missions. During ground hold (i.e., prior to launch) moisture condensation must be minimized in the gap between the cryo-insulation and the TPS. The longer term solution requires the development of a single material to act as cryo- insulation during ground hold and as TPS during re-entry. Such a material minimizes complexity and weight while improving reliability and reducing cost.

Lambert, Michael A.↗
NASA NTRS2017

Thin Film Thermal Conductivity Measurements using Superconducting Nanowires

We present a simple experimental scheme for estimating the cryogenic thermal transport properties of thin films using superconducting nanowires. In a parallel array of nanowires, the heat from one nanowire in the normal state changes the local temperature around adjacent nanowires, reducing their switching current. Calibration of this change in switching current as a function of bath temperature provides an estimate of the temperature as a function of displacement from the heater. This provides a method of determining the contribution of substrate heat transport to the cooling time of superconducting nanowire single photon detectors. Understanding this process is necessary for successful electrothermal modeling of superconducting nanowire systems.

Shaw, M. D.↗
NASA NTRS2013

Development and Testing of a Variable Conductance Thermal Acquisition, Transport, and Switching System

This paper describes the development and testing of a scalable thermal control architecture for instruments, subsystems, or systems that must operate in severe space environments with wide variations in sink temperature. The architecture is comprised by linking one or more hot-side variable conductance heat pipes (VCHPs) in series with one or more cold-side loop heat pipes (LHPs). The VCHPs provide wide area heat acquisition, limited distance thermal transport, modest against gravity pumping, concentrated LHP startup heating, and high switching ratio variable conductance operation. The LHPs provide localized heat acquisition, long distance thermal transport, significant against gravity pumping, and high switching ratio variable conductance operation. Combining two variable conductance devices in series ensures very high switching ratio isolation from severe environments like the Earth's moon, where each lunar day spans 15 Earth days (270 K sink, with a surface-shielded/space viewing radiator) and each lunar night spans 15 Earth days (80-100 K radiative sink, depending on location). The single VCHP-single LHP system described herein was developed to maintain thermal control of International Lunar Network (ILN) anchor node lander electronics, but it is also applicable to other variable heat rejection space missions in severe environments. The LHPVCHP system utilizes a stainless steel wire mesh wick ammonia VCHP, a Teflon wick propylene LHP, a pair of one-third square meter high  radiators (one capillary-pumped horizontal radiator and a second gravity-fed vertical radiator), a half-meter of transport distance, and a wick-bearing co-located flow regulator (CLFR) to allow operation with a hot (deactivated) radiator. The VCHP was designed with a small reservoir formed by extending the length of its stainless steel heat pipe tubing. The system was able to provide end-to-end switching ratios of 300-500 during thermal vacuum testing at ATK, including 3-5 W/K ON conductance and 0.01 W/K OFF conductance. The test results described herein also include an in-depth analysis of VCHP condenser performance to explain VCHP switching operation in detail. Future multi-VCHP/multi-LHP thermal management system concepts that provide scalability to higher powers/longer transport lengths are also discussed in the paper.

Bugby, D. C.↗
NASA NTRS2013

Development and Testing of a Variable Conductance Thermal Acquisition, Transport, and Switching System

This paper describes the development and testing of a scalable thermal management architecture for instruments, subsystems, or systems that must operate in severe space environments with wide variations in sink temperature. The architecture involves a serial linkage of one or more hot-side variable conductance heat pipes (VCHPs) to one or more cold-side loop heat pipes (LHPs). The VCHPs provide wide area heat acquisition, limited distance thermal transport, modest against gravity pumping, concentrated LHP startup heating, and high switching ratio variable conductance operation. The LHPs provide localized heat acquisition, long distance thermal transport, significant against gravity pumping, and high switching ratio variable conductance operation. The single-VCHP, single-LHP system described herein was developed to maintain thermal control of a small robotic lunar lander throughout the lunar day-night thermal cycle. It is also applicable to other variable heat rejection space missions in severe environments. Operationally, despite a 60-70% gas blocked VCHP condenser during ON testing, the system was still able to provide 2-4 W/K ON conductance, 0.01 W/K OFF conductance, and an end-to-end switching ratio of 200-400. The paper provides a detailed analysis of VCHP condenser performance, which quantified the gas blockage situation. Future multi-VCHP/multi-LHP thermal management system concepts that provide power/transport length scalability are also discussed.

Bugby, David C.↗
NASA NTRS1993

High thermal conductivity of diamond

The objectives of this educational exercise were to demonstrate the high rate of heat flow from a synthetic diamond coupon and to compare it to a commonly used thermal conductor, such as copper. The principles of heat transfer by conduction and convection may also be demonstrated. A list of equipment and supplies and the procedure for the experiment are presented.

Stephan, Patrick M.↗
NASA NTRS1997

Development of Tailorable Electrically Conductive Thermal Control Material Systems

The optical characteristics of surfaces on spacecraft are fundamental parameters in controlling its temperature. Passive thermal control coatings with designed solar absorptance and infrared emittance properties have been developed and have been in use for some time. In this total space environment, the coating must be stable and maintain its desired optical properties as well as mechanical properties for the course of the mission lifetime. The mission lifetimes are increasing and in our quest to save weight, newer substrates are being integrated which limit electrical grounding schemes. All of this has added to already existing concerns about spacecraft charging and related spacecraft failures or operational failures. The concern is even greater for thermal control surfaces that are very large. One way of alleviating such concerns is to design new thermal control material systems (TCMS) that can help to mitigate charging via providing charge leakage paths. The objective of this program was to develop two types of passive electrically conductive TCMS. The first was a highly absorbing/emitting black surface and the second was a low (alpha(sub s)/epsilon(sub N)) type white surface. The surface resistance goals for the black absorber was 10(exp 4) to 10(exp 9) Omega/square, and for the white surfaces it was 10(exp 6) to 10(exp 10) Omega/square. Several material system concepts were suggested and evaluated for space environment stability and electrical performance characterization. Our efforts in designing and evaluating these material systems have resulted in several developments. New concepts, pigments and binders have been developed to provide new engineering quality TCMS. Some of these have already found application on space hardware, some are waiting to be recognized by thermal designers, and some require further detailed studies to become state-of-the-art for future space hardware and space structures. Our studies on baseline state-of-the-art materials and conductive concepts have resulted in several important findings that are of interest to all thermal designers and systems integrators.

Deshpande, M. S.↗
NASA NTRS1990

Global effects of thermal conduction on two-phase media

The evolution of two-phase systems of astrophysical gases which change mass between the phases is studied to see whether a steady state is ever reached. The criterion for thermal instability in a cloudy medium is derived. The evolution of the pressure and density of the intercloud medium under the combined effects of heating and radiative cooling of the intercloud gas on the one hand and evaporation and condensation of the clouds on the other is determined. The equilibrium density to which the system evolves is determined for the case when the pressure is fixed. The theory is illustrated by the case in which the intercloud gas is heated by Compton scattering in a hard radiation field and cooled by bremsstrahlung and inverse Compton scattering.

Begelman, Mitchell C.↗
DOE OSTI2020

Packaging Development for a 1200V SiC BiDFET Switch Using Highly Thermally Conductive Organic Epoxy Laminate

A novel 1.2 kV/10A, 4H-SiC monolithic, bidirectional switch has been developed for use in cycloconverter applications to facilitate high-frequency direct AC-to-AC power conversion and enables new power converter topologies. A new packaging solution, utilizing a 100 µm flexible polyimide organic laminate substrate is developed to mitigate thermo-mechanical stress during power cycling and enable smaller form factor and lower cost. Multiphysics simulations and static tests were conducted to show performance characterization of the module and compare it against metallic substrates. A new organic laminate epoxy resin composite dielectric (ERCD) is also evaluated for superior thermal performance and shows 63% reduction in junction to case resistance compared to DBC substrates.

Silicon Carbide, Bi-directional switch, BiDFET, MO↗
NASA NTRS2007

Comparison of GRCop-84 to Other High Thermal Conductive Cu Alloys

The mechanical properties of five copper alloys (GRCop-84, AMZIRC, GlidCop Al-15, Cu-1Cr-0.1Zr, Cu-0.9Cr) competing in high temperature, high heat flux applications such as rocket nozzles, were compared. Tensile, creep, thermal expansion, and compression tests are presented. Tests were done on as-received material, and on material which received a simulated brazing heat treatment at 935 C. The 935 C heat treatment weakened AMZIRC, Cu-1Cr-0.1Zr, and Cu-0.9Cr, and the strength of as-received AMZIRC dropped precipitously as test temperatures exceeded 500 C. The properties of GlidCop Al-15 and GRCop-84 were not significantly affected by the 935 C heat treatment. Thus GRCop-84 is better than AMZIRC, Cu-1Cr-0.1Zr, and Cu-0.9Cr at temperatures greater than 500 C. Ductility was lowest in GlidCop Al-15 and Cu-0.9Cr. The creep properties of GRCop-84 were superior to those of brazed AMZIRC, Cu-1Cr-0.1Zr, and Cu-0.9Cr. At equivalent rupture life and stress, GRCop-84 had a 150 C temperature advantage over brazed AMZIRC; for equivalent rupture life and temperature GRCop-84 was two times stronger. The advantages of GRCop-84 over GlidCop Al-15 associated with ease of processing were confirmed by GlidCop s marginal ductility. In the post brazed condition, GRCop-84 was found to be superior to the other alloys due to its greater strength and creep resistance (compared to AMZIRC, Cu-1Cr-0.1Zr, and Cu-0.9Cr) and ductility (compared to GlidCop Al-15)

deGroh, Henry C., III↗
NASA NTRS1998

Development of Tailorable Electrically Conductive Thermal Control Material Systems

The optical characteristics of surfaces on spacecraft are fundamental parameters in controlling its temperature. Passive thermal control coatings with designed solar absorptance and infrared emittance properties have been developed and been in use for some time. In this total space environment, the coating must be stable and maintain its desired optical properties for the course of the mission lifetime. The mission lifetimes are increasing and in our quest to save weight, newer substrates are being integrated which limit electrical grounding schemes. All of this has already added to the existing concerns about spacecraft charging and related spacecraft failures or operational failures. The concern is even greater for thermal control surfaces that are very large. One way of alleviating such concerns is to design new thermal control material systems (TCMS) that can help to mitigate charging via providing charge leakage paths. The object of this program was to develop two types of passive electrically conductive TCMS.

Deshpande, M. S.↗
NASA NTRS2016

X-Ray Source Heights in a Solar Flare: Thick-Target Versus Thermal Conduction Front Heating

Observations of solar flares with RHESSI have shown X-ray sources traveling along flaring loops, from the corona down to the chromosphere and back up. The 2002 November 28 C1.1 flare, first observed with RHESSI by Sui et al. and quantitatively analyzed by O'Flannagain et al., very clearly shows this behavior. By employing numerical experiments, we use these observations of X-ray source height motions as a constraint to distinguish between heating due to a non-thermal electron beam and in situ energy deposition in the corona. We find that both heating scenarios can reproduce the observed light curves, but our results favor non-thermal heating. In situ heating is inconsistent with the observed X-ray source morphology and always gives a height dispersion with photon energy opposite to what is observed.

Reep, J. W.↗
DOE OSTI2024

Thermal Conductivity Measurements on Irradiated TRISO Fuel at IMCL

INL is leading the development and testing of accident tolerant fuels: fuels that resist melting down and that contain radioactive byproducts, preventing their release into their environment. One such fuel, called TRi-structural ISOtropic particle fuel (TRISO) See Fig 1. Understanding how heat propagates through a fuel is essential to fuel development and qualification. When irradiated in a reactor the properties that affect heat transport change. Our objective is to quantify the change in thermal properties of TRISO after irradiation.

11 NUCLEAR FUEL CYCLE AND FUEL MATERIALS↗