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At least 829 records · Page 46

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NASA NTRS1987

Thermal conductivity of pristine and brominated highly graphitized pitch based carbon fibers

Thermal conductivity of brominated and pristine Union Carbide P-100 graphite fibers in the 30 to 160 C temperature range was determined by measuring thermal conductivities of graphite fiber epoxy composite samples and then excluding the epoxy contribution. A comparative thermal conductivity instrument was used to measure the thermal conductivity of the samples containing fibers. Results showed that the thermal conductivity values were 225 to 370 w/m-K and 215 to 340 w/m-K for pristine and brominated fibers, respectively. Furthermore, the thermal conductivity ratio of brominated to pristine P-100 fibers was 0.89, 0.91, and 0.92 at 55 to 80 C; 108 and 130 C, respectively. Such decrease in thermal conductivity results almost entirely from the 10 percent increase in fiber cross sectional area due to bromination. This result suggests that bromination effects on P-100 fiber structure is small, and that some structural changes, presumably the sharp-angled domain wall becomes less sharp, occurring in the 80 to 108 C temperature range.

Hung, Ching-Cheh↗

NASA NTRS1988

Thermal cycling of tungsten-fibre-reinforced superalloy composites

The thermal cycling of a tungsten-fiber-reinforced superalloy (TFRS) composite is typical of its application in high-temperature engine environments. The mismatch in thermal expansion coefficients between fiber and matrix causes substantial longitudinal (0 deg) stresses in the composite, which can produce inelastic damage-producing matrix strains. The case of thermal fatigue is explored as a "worst case" of the possible matrix damage, in comparison with specimens which are also mechanically loaded in tension. The thermally generated cyclic stresses and the attendant matrix plasticity may be estimated using a nonlinear finite-element program, by proposing a physical analog to the micromechanics equations. A damage metric for the matrix is proposed using the Coffin-Manson criterion, which metric can facilitate comparisons of damage among different candidate materials, and also comparisons for a given material subjected to different temperature cycles. An experimental program was carried out for thermal cycling of a 37 vol pct TFRS composite to different maximum temperatures. The results confirm the prediction that thermal cycling produces matrix degradation and composite strength reduction, which become more pronounced with increasing maximum cyclic temperature. The strength of the fiber is shown to be identical for the as-fabricated and thermally cycled specimens, suggesting that the reduction in composite strength is due to the loss of matrix contribution and also to notching effects of the matrix voids on the fiber.

Wetherhold, Robert C.↗

NASA NTRS1988

Materials selection for long life in LEO: A critical evaluation of atomic oxygen testing with thermal atom systems

The use of thermal atom test methods as a materials selection and screening technique for low-Earth orbit (LEO) spacecraft is critically evaluated. The chemistry and physics of thermal atom environments are compared with the LEO environment. The relative reactivities of a number of materials determined to be in thermal atom environments are compared to those observed in LEO and in high quality LEO simulations. Reaction efficiencies measured in a new type of thermal atom apparatus are one-hundredth to one-thousandth those observed in LEO, and many materials showing nearly identical reactivities in LEO show relative reactivities differing by as much as a factor of 8 in thermal atom systems. A simple phenomenological kinetic model for the reaction of oxygen atoms with organic materials can be used to explain the differences in reactivity in different environments. Certain specific thermal test environments can be used as reliable materials screening tools. Using thermal atom methods to predict material lifetime in LEO requires direct calibration of the method against LEO data or high quality simulation data for each material.

Koontz, S. L.↗

NASA NTRS1988

Thermal distortion analysis of the Space Station solar dynamic concentrator

A method was developed to evaluate the thermal distortion of the Space Station Solar Dynamic Concentrator and the effects of thermal distortion on concentrator optical performance. The analytical method includes generating temperature distributions with TRASYS and SINDA models, interfacing the SINDA results with the SINDA-NASTRAN Interface Program (SNIP), calculating thermal distortion with a NASTRAN/PATRAN finite element model, and providing flux distribution maps within the receiver with the ray tracing OFFSET program. Temperature distributions, thermally induced slope errors, and flux distribution maps within the receiver are discussed. Results during a typical orbit indicate that temperatures of the hexagonal panels and triangular facets range between -18 and 99 C (-1 to 210 F), facet rotations are less than 0.2 mrad, and a change in facet radius due to thermal flattening is less than 5 percent. The predicted power loss with thermal distortion effects was less than 0.3 percent. The thermal distortion of the Solar Dynamic concentrator has negligible effect on the flux distribution within the receiver cavity.

Trudell, Jeffery J.↗

NASA NTRS1988

Durability of thermal barrier coatings in a high heat flux environment

Thermal shock is a significant factor in the limited service life of Space Shuttle Main Engine (SSME) high pressure, fuel turbopump (HPFTP) turbine blades. Addition of advanced thermal barrier coatings (TBCs) to the blades could serve to dampen the thermal shock, thereby increasing the life of the blades. However, testing and use of TBCs to date is performed primarily under moderate heat flux conditions which are typical of aircraft turbines. Only limited testing was conducted that addresses high heat flux and severe thermal shock conditions. Therefore, it is not clear if TBCs can survive severe thermal shocks or provide adequate thermal shock protection to the HPFTP turbine blades. The purpose is to experimentally evaluate the potential durability and protective capability of a variety of advanced TBCs in a cyclic thermal shock environment. A secondary goal is to identify significant parameters affecting TBC life during high heat flux testing. Parameters investigated include top coat thickness, bond coat thickness, substrate type, and substrate geometry.

Brindley, William J.↗

NASA NTRS1990

Thermal diffusivity of diamond films using a laser pulse technique

Polycrystalline diamond films were deposited using a microwave plasma-enhanced chemical vapor deposition process. A laser pulse technique was developed to measure the thermal diffusivity of diamond films deposited on a silicon substrate. The effective thermal diffusivity of a diamond film on silicon was measured by observing the phase and amplitude of the cyclic thermal waves generated by laser pulses. An analytical model is presented to calculate the effective in-plane (face-parallel) diffusivity of a two-layer system. The model is used to reduce the effective thermal diffusivity of the diamonds/silicon sample to a value for the thermal diffusivity and conductivity of the diamond film. The average effective diffusivity values are 1.47 + or - 0.03 and 1.83 + or - 0.10 yielding thermal diffusivity values of 7.46 + or - 0.90 and 7.33 + or - 0.70 sq cm/s respectively, for the two samples; the calculated thermal con ductivity values are 13.50 and 13.28 W/cmK, which are better than that of type 1a natural diamond. The phase and amplitude measurements give similar results.

Albin, Sacharia↗

NASA NTRS1990

Thermal grease replacement for the modular power subsystem

A comparative thermal conductance test was conducted to evaluate thermal interface materials for use on the Modular Power Subsystem. Materials tested included three thermal pads, four RTV adhesives, bare metal, and one thermal grease. The tests were conducted in a bell jar at vacuum conditions using a 1400 square centimeter footprint and two relatively low contact pressures, 207 kPa and 620 kPa. Power inputs ranged from 100 to 500 watts, and the thermal interface conductance values ranged from 100 to 1700 W/m2 C for the interstitial materials tested. In general, the thermal pads performed a little better than bare metal, while the RTV adhesives performed significantly better than the bare metal and comparable to the thermal grease.

Lapinski, John R., Jr.↗

NASA NTRS1990

Finite element analysis of thermal distortion effects on optical performance of solar dynamic concentrator for Space Station Freedom

An analysis was performed to predict the thermal distortion of the solar dynamic concentrator for Space Station Freedom in low earth orbit and to evaluate the effects of that thermal distortion on concentrator on-orbit performance. The analysis required substructural finite element modeling of critical concentrator structural subsystems, structural finite element modeling of the concentrator, mapping of thermal loading onto the structural finite element model, and the creation of specialized postprocessors to assist in interpreting results. Concentrator temperature distributions and thermally induced displacements and slope errors and the resulting receiver flux distribution profiles are discussed. Results determined for a typical orbit indicate that concentrator facet rotations are less than 0.2 mrad and that the change in facet radius due to thermal flattening is less than 5 percent. The predicted power loss due to thermal distortion effects is less than 0.3 percent. As a consequence the thermal distortions of the solar dynamic concentrator in low earth orbit will have a negligible effect on the flux distribution profiles within the receiver.

Doherty, Michael P.↗

NASA NTRS1990

Thermal control for the ground simulation of a space IR sensor system

Active and passive thermal controls for simulations of a space IR sensor system operating in cryogenic temperatures were designed, built, and tested from a component level to a system level. The test results from component tests and integrated system tests have compared very well with theoretical predictions, and thus verify component and integrated thermal math models. These verified models can be modified for use to predict flight systems thermal performance. Thermal vacuum simulations and demonstrations of a space IR system consisted of a target and background scene generator, telescope mirrors supported by a graphite-epoxy metering structure, and an IR sensor. These components are required to operate at cryogenic temperature levels. Each component has its unique thermal control needs. Descriptions are presented of thermal control systems for the test article from component design level to integrated system level along with discussions of component and integrated demonstration tests, and correlation of test data with thermal finite difference models.

Lee, E.↗

NASA NTRS1990

Space environmental effects on spacecraft thermal control coatings

A large portion of the space based applications in the near future are for low Earth orbit (LEO) including the Space Station Freedom (SSF). The lifetime needs for materials on the SSF is 30 years. Materials must be selected which can withstand the deleterious effects of LEO. Environmental effects in the LEO include atomic oxygen, UV, ionizing radiation, and hypervelocity impact. These effects can adversely affect the surface properties of materials. This is particularly critical in the case of thermal control materials where the efficiency of the thermal control is dependent on the stability of the surface properties. The current baseline thermal coating for the SSF radiators is Ag Teflon. The surface property requirements for the coatings are a solar absorptance of 0.2 and an infrared emittance of 0.8. The effects of atomic oxygen and UV radiation on the baseline coating and several other candidate thermal control materials were studied. The thermal control radiator materials included Ag and Al backed Teflon, H2SO4 anodized Al, sputter deposited SiO2 on Al, and Ag and Al backed polychloro trifluoroethylene. The simulation of several of the LEO environment constituents provided a data base to aid in the selection of the radiator thermal control material to meet the life requirements of the SSF. The effects are illustrated of the environment on thermal control coatings and the importance of this factor in the selection process for long life spacecraft materials.

Bartosik, L. G.↗

NASA NTRS1991

Thermal shock testing for assuring reliability of glass-sealed microelectronic packages

Tests were performed to determine if thermal shocking is destructive to glass-to-metal seal microelectronic packages and if thermal shock step stressing can compare package reliabilities. Thermal shocking was shown to be not destructive to highly reliable glass seals. Pin-pull tests used to compare the interfacial pin glass strengths showed no differences between thermal shocked and not-thermal shocked headers. A 'critical stress resistance temperature' was not exhibited by the 14 pin Dual In-line Package (DIP) headers evaluated. Headers manufactured in cryogenic nitrogen based and exothermically generated atmospheres showed differences in as-received leak rates, residual oxide depths and pin glass interfacial strengths; these were caused by the different manufacturing methods, in particular, by the chemically etched pins used by one manufacturer. Both header types passed thermal shock tests to temperature differentials of 646 C. The sensitivity of helium leak rate measurements was improved up to 70 percent by baking headers for two hours at 200 C after thermal shocking.

Thomas, Walter B., III↗

NASA NTRS1991

Thermal analysis of conceptual designs for GPHS/FPSE power systems of 250 We and 500 We

Thermal analyses were performed for two distinct configurations of a proposed space nuclear power system which combines General Purpose Heat Source (GPHS) modules with the state of the art Free-Piston Stirling Engines (FPSEs). The two configurations correspond to systems with power levels of 250 and 500 W(sub e). The 250 W(sub e) GPHS/FPSE power system utilizes four GPHS modules and one FPSE, and the 500 W(sub e) contains eight GPHS modules and two FPSEs. The configurations of the systems and the bases for selecting the configurations are described. Brief introductory sections are included to describe the GPHS modules and free piston Stirling engines. The primary focus of the thermal analyses is on the temperature of the iridium fuel clad within the GPHS modules. A design goal temperature of 1573 K was selected as the upper limit for the fuel clad during normal operating conditions. The basis for selecting this temperature limit is discussed in detail. Results obtained from thermal analysis of the 250 W(sub e) GPHS/FPSE power system indicate fuel clad temperatures which slightly exceed the design goal temperature of 1573 K. The results are considered favorable due to the numerous conservative assumptions used in developing the thermal model and performing the thermal analysis. To demonstrate the effects of the conservatism, a brief sensitivity analysis is performed in which a few of the key system parameters are varied to determine their effect on the fuel clad temperatures. It is concluded that thermal analysis of a more detailed thermal model would be expected to yield fuel clad temperatures below the design foal temperature limiy 1573 K.

Mccomas, Thomas J.↗

NASA NTRS1991

Materials and light thermal structures research for advanced space exploration

The Light Thermal Structures Center at the University of Virginia sponsors educational and research programs focused on the development of reliable, lightweight structures to function in hostile thermal environments. Technology advances in materials and design methodology for light thermal structures will contribute to improved space vehicle design concepts with attendant weight savings. This paper highlights current research activities in three areas relevant to space exploration: low density, high temperature aluminum alloys, composite materials, and structures with thermal gradients. Advances in the development of new aluminum-lithium alloys and mechanically alloyed aluminum alloys are described. Material properties and design features of advanced composites are highlighted. Research studies in thermal structures with temperature gradients include inelastic panel buckling and thermally induced unstable oscillations. Current and future research is focused on the integration of new materials with applications to structural components with thermal gradients.

Thornton, Earl A.↗

NASA NTRS1991

Thermal response of large area high temperature superconducting YBaCuO infrared bolometers

Thermal analysis of large area high temperature superconducting infrared detector operating in the equilibrium mode (bolometer) was performed. An expression for the temperature coefficient beta = 1/R(dR/dT) in terms of the thermal conductance and the thermal time constant of the detector were derived. A superconducting transition edge bolometer is a thermistor consisting of a thin film superconducting YBaCuO evaporated into a suitable thermally isolated substrate. The operating temperature of the bolometer is maintained close to the midpoint of the superconducting transition region where the resistance R has a maximum dynamic range. A detector with a strip configuration was analyzed and an expression for the temperature rise (delta T) above the ambient due to a uniform illumination with a source of power density was calculated. An expression for the thermal responsibility depends upon the spatial modulation frequency and the angular frequency of the incoming radiation. The problem of the thermal cross talk between different detector elements was addressed. In the case of monolithic HTS detector array with a row of square elements of dimensions 2a and CCD or CID readout electronics the thermal spread function was derived for different spacing between elements.

Khalil, Ali E.↗

NASA NTRS1992

Coupling device with improved thermal interface

The primary object of the present invention is to provide a simple, reliable, and lightweight coupling that will also have an efficient thermal interface. A further object of the invention is to provide a coupling that is capable of blind mating with little or no insertion forces. Another object of the invention is to provide a coupling that acts as a thermal regulator to maintain a constant temperature on one side of the coupling. Another object of the invention is to increase the available surface area of a coupling thus providing a larger area for the conduction of heat across the thermal interface. Another object of the invention is to provide a fluidic coupling that has no fluid passing across the interface, thus reducing the likelihood of leaks and contamination. The foregoing objects are achieved by utilizing, as in the prior art, a hot area (at an elevated temperature as compared to a cold area) with a need to remove excess heat from the hot area to a cold area. In this device, the thermal interface will occur not on a planar horizontal surface, but along a non-planar vertical surface, which will reduce the reaction forces and increase the thermal conductivity of the device. One non-planar surface is a surface on a cold pin extending from the cold area and the other non-planar surface is a surface on a hot pin extending from the hot area. The cold pin is fixed and does not move while the hot pin is a flexible member and its movement towards the cold pin will bring the two non-planar surfaces together forming the thermal interface. The actuating member for the device is a shape-memory actuation wire which is attached through an aperture to the hot pin and through another aperture to an actuation wire retainer. By properly programming the actuation wire, heat from the hot area will cause the actuation wire to bend the hot wire. Heat from the hot area will cause the actuation wire to bend the hot pin towards the cold pin forming the coupling and the desired thermal interface. The shape-memory actuation wire is made of a shape-memory-effect alloy such as Nitinol.

Milam, Malcolm Bruce↗

NASA NTRS1992

Simulation of the synergistic low Earth orbit effects of vacuum thermal cycling, vacuum UV radiation, and atomic oxygen

In order to assess the low Earth orbit (LEO) durability of candidate space materials, it is necessary to use ground laboratory facilities which provide LEO environmental effects. A facility combining vacuum thermal cycling and vacuum ultraviolet (VUV) radiation has been designed and constructed at NASA Lewis Research Center for this purpose. This facility can also be operated without the VUV lamps. An additional facility can be used to provide VUV exposure only. By utilizing these facilities, followed by atomic oxygen exposure in an RF plasma asher, the effects of the individual vacuum thermal cycling and VUV environments can be compared to the effect of the combined vacuum thermal cycling/VUV environment on the atomic oxygen durability of materials. The synergistic effects of simulated LEO environmental conditions on materials were evaluated by first exposing materials to vacuum thermal cycling, VUV, and vacuum thermal cycling/VUV environments followed by exposure to atomic oxygen in an RP plasma asher. Candidate space power materials such as atomic oxygen protected polyimides and solar concentrator mirrors were evaluated using these facilities. Characteristics of the Vacuum Thermal Cycling/VUV Exposure Facility which simulates the temperature sequences and solar ultraviolet radiation exposure that would be experienced by a spacecraft surface in LEO are discussed. Results of durability evaluations of some candidate space power materials to the simulated LEO environmental conditions will also be discussed. Such results have indicated that for some materials, atomic oxygen durability is affected by previous exposure to thermal cycling and/or VUV exposure.

Dever, Joyce A.↗

NASA NTRS1993

Thermal expansion behavior of LDEF metal matrix composites

The thermal expansion behavior of Long Duration Exposure Facility (LDEF) metal matrix composite materials was studied by (1) analyzing the flight data that was recorded on orbit to determine the effects of orbital time and heating/cooling rates on the performance of the composite materials, and (2) characterizing and comparing the thermal expansion behavior of post-flight LDEF and lab-control samples. The flight data revealed that structures in space are subjected to nonuniform temperature distributions, and thermal conductivity of a material is an important factor in establishing a uniform temperature distribution and avoiding thermal distortion. The flight and laboratory data showed that both Gr/Al and Gr/Mg composites were stabilized after prolonged thermal cycling on orbit. However, Gr/Al composites showed more stable thermal expansion behavior than Gr/Mg composites and offer advantages for space structures particularly where very tight thermal stability requirements in addition to high material performance must be met.

Le, Tuyen D.↗

NASA NTRS1993

Space Station thermal storage/refrigeration system research and development

Space Station thermal loading conditions represent an order of magnitude increase over current and previous spacecraft such as Skylab, Apollo, Pegasus III, Lunar Rover Vehicle, and Lockheed TRIDENT missiles. Thermal storage units (TSU's) were successfully used on these as well as many applications for ground based solar energy storage applications. It is desirable to store thermal energy during peak loading conditions as an alternative to providing increased radiator surface area which adds to the weight of the system. Basically, TSU's store heat by melting a phase change material (PCM) such as a paraffin. The physical property data for the PCM's used in the design of these TSU's is well defined in the literature. Design techniques are generally well established for the TSU's. However, the Space Station provides a new challenge in the application of these data and techniques because of three factors: the large size of the TSU required, the integration of the TSU for the Space Station thermal management concept with its diverse opportunities for storage application, and the TSU's interface with a two-phase (liquid/vapor) thermal bus/central heat rejection system. The objective in the thermal storage research and development task was to design, fabricate, and test a demonstration unit. One test article was to be a passive thermal storage unit capable of storing frozen food at -20 F for a minimum of 90 days. A second unit was to be capable of storing frozen biological samples at -94 F, again for a minimum of 90 days. The articles developed were compatible with shuttle mission conditions, including safety and handling by astronauts. Further, storage rack concepts were presented so that these units can be integrated into Space Station logistics module storage racks. The extreme sensitivity of spacecraft radiator systems design-to-heat rejection temperature requirements is well known. A large radiator area penalty is incurred if low temperatures are accommodated via a single centralized radiator system. As per the scope of work of this task, the applicability of refrigeration system tailored to meet the specialized requirements of storage of food and biological samples was investigated. The issues addressed were the anticipated power consumption and feasible designs and cycles for meeting specific storage requirements. Further, development issues were assessed related to the operation of vapor compression systems in micro-gravity addressing separation of vapor and liquid phases (via capillary systems).

Dean, W. G.↗