Titles:

• A Review of Textured Surfaces, Paints, and Coatings for Space Radiator Applications   February 2004
  

• Thermal Performance of an Annealed Pyrolytic Graphite Solar Collector  February 2002
   
• Electrical and Thermal Conductivity of Carbon Fiber-Polymer Composite Plates  2002

• Modeling Heat Flow in a Calorimeter Equipped With a Textured Solar Collector  August 2001

• Thermal Emittance Measurements on Candidate Refractory Materials for Application in Nuclear Space Power Systems  February 2001
   
• Portable Infrared Reflectometer for Evaluating Emittance  February 2000
   
• Optical Property Enhancement and Durability Evaluation of Heat Receiver Aperture Shield Materials  January 1998
   
• Correlation of Predicted and Observed Optical Properties of Multi-Layer Thermal Control Coatings  1998
   
• Effects of Ambient High Temperature Exposure on Alumina-Titania High Emittance Surfaces for Solar Dynamic Systems  1998
   
• Reflectivity of Silver and Silver-coated Substrates from 25°C to 800°C  July 1997
   
• Advances in Optical Property Measurements of Spacecraft Materials  June 1997
   
• Optical and Calorimetric Evaluation of Z-93-P and Other Thermal Control Coatings   1996
   
• Emittance of Thermal Control Materials Between 100K and 400K  November 1996

• Emissivity Characterization of Plastics, Ceramics, and Coatings Using a Calorimetric Technique  May 1995
   
• Performance and Durability of High Emittance Heat Receiver Surfaces for Solar Dynamic Power Systems  March 1994
   
• Emittance Characterization of Thermal Control Paints, Coatings, and Surfaces Using a Calorimetric Technique  1994
   
• Thermal Modeling of a Calorimetric Technique for Measuring the Emittance of Surfaces and Coatings  1993

Jaworske, D. A., “A Review of Textured Surfaces, Paints, and Coatings for Space Radiator Applications,” Space Technology & Applications International Forum (STAIF-2004), Albuquerque, NM, pp. 71-75, February 2004.

Future space nuclear power systems will require large radiators to dissipate excess thermal energy.  Such radiators may be composed of carbon-carbon composite fins made from high thermal conductivity graphite fibers or may be a more traditional honeycomb structure with face sheets composed of a suitable high temperature metal.  In either case, the surface of the radiator must have a high emittance at the desired operating temperature, envisioned to be in the range of 400 to 900 K, and must be durable for the length of the mission, envisioned to be ten years.  Existing thermal control paints and coatings may be applicable at the low end of the envisioned temperature range, but may not be applicable at elevated temperatures.  Hence, other avenues of emittance enhancement need to be explored.  Previous work has identified a number of promising technologies that may be useful for enhancing the emittance of candidate surfaces, including texturing the radiator surface via sand blasting, oxidation at elevated temperature, and exposure to atomic oxygen.  This paper will review existing candidate thermal control paints and coatings to identify their strengths and weaknesses and will review other promising technologies that have been proposed in the past few years to enhance the emittance of radiator surfaces.

A solar collector having the combined properties of high solar absorptance, low infrared emittance, and high thermal conductivity is needed for applications where solar energy is to be absorbed and transported for use in minisatellites. Such a solar collector may be used with a low temperature differential heat engine to provide power or with a thermal bus for thermal switching applications. One concept being considered for the solar collector is an Al-Al2O3 cermet coating applied to a thermal conductivity enhanced polished aluminum substrate. The cermet coating provides high solar absorptance, and the polished aluminum provides low infrared emittance. Annealed pyrolytic graphite embedded in the aluminum substrate provides enhanced thermal conductivity. The as-measured thermal performance of an annealed pyrolytic graphite thermal conductivity enhanced polished aluminum solar collector, coated with a cermet coating, will be presented.
 

Gaier, J. R., Stueben, H., Berkebile, S., and Balagadde, F., “Electrical and Thermal Conductivity of Carbon Fiber-Polymer Composite Plates,” abstracted in Ninth International Conference on Composite Engineering, (D. Hui Ed.), International Community for Composites Engineering, pp 217-8, San Diego, 2002.

Carbon fiber-polymer composite plates were fabricated using 0°-90° woven fabrics of a variety of pristine and bromine intercalated carbon fibers. The fibers had electrical resistivities varying from 50 to 1800 µ ohm-cm, and thermal conductivities varying from 8.5 to 520 W/m-K. Anisotropic composites were also fabricated from fabrics with low conductivity fibers in the warp direction and high conductivity in the weft. Composite electrical resistivity was measured using an eddy current technique and a four-point technique, and calculated using a geometry- corrected rule of mixtures. Composite thermal conductivity was measured using an optical heating technique and infrared scanning of the surface as well as being calculated from the rule of mixtures. Woven fabrics were shown to behave like homogeneous, isotropic plates both electrically and thermally as long as the samples are large with respect to the weave size of the fabric. The four-point resistivity was somewhat higher than that predicted by the rule of mixtures. The resistivity as measured by the eddy current method was in all cases higher than both the four-point and rule of mixture resistivities. The thermal conductivities of the composite were in fairly good agreement with the rule of mixtures for relatively low conductivity fibers, but much lower than predicted for high conductivity fibers. Anisotropic composites could only be made by stacking the anisotropic fabrics in a 0°-0° geometry. Even under those conditions the anisotropy, especially of the thermal conductivity, was considerably less than would be expected from the rule of mixtures.

Jaworske, D. A. and Allen, B. J., “Modeling Heat Flow in a Calorimeter Equipped With a Textured Solar Collector,” 36th IECEC, Savannah, GA, pp. 1195-1200, August 2001. 

Heat engines are being considered for generating electric power for minisatellite applications, particularly for those missions in high radiation threat orbits.  To achieve this objective, solar energy must be collected and transported to the hot side of the heat engine.  A solar collector is needed having the combined properties of high solar absorptance, low infrared emittance, and high thermal conductivity.  To test candidate solar collector concepts, a simple calorimeter was designed, manufactured, and installed in a bench top vacuum chamber to measure heat flow.  In addition, a finite element analysis model of the collector/calorimeter combination was made to model this heat flow.  The model was tuned based on observations from the as-manufactured collector/calorimeter combination.  In addition, the model was exercised to examine other collector concepts, properties, and scale up issues

The development of a highly efficient General Purpose Heat Source (GPHS) space power system requires that all of the available thermal energy from the GPHS modules be utilized in the most thermally efficient manner. This includes defining the heat transfer/thermal gradient profile between the surface of the GPHS’s and the surface of the energy converter’s hot end whose geometry is dependent on the converter technology selected. Control of the radiant heat transfer between these two surfaces is done by regulating how efficiently the selected converter’s hot end surface can reject heat compared to a perfect blackbody, i.e. its infrared emittance. Several refractory materials of interest including niobium-1% zirconium, molybdenum-44.5% rhenium and L-605 (a cobalt-based alloy) were subjected to various surface treatments (grit blasting with either SiC or WC particulates) and heat treatments (up to 1198 K for up to 3000 hours). Room temperature infrared emittance values were then obtained using two different infrared reflectometers. Grit blasting with either SiC or WC tended to increase the emittance of flat or curved L-605 coupons by ~0.2-0.3 independent of heat treatment. Heat treating L-605 coupons under 773 K for up to 2000 hours had only a slight effect on their emittance, while heat treating L-605 coupons at 973 K for over 150 hours appeared to significantly increase their emittance. For the temperatures and times studied here, the emittance values obtained on niobium-1% zirconium and molybdenum-44.5% rhenium coupons were independent of heat treat times and temperatures (except for the niobium-1% zirconium coupon that was heat treated at 1198 K for 150 hours).

Optical methods are frequently used to evaluate the emittance of candidate spacecraft thermal control materials. One new optical method utilizes a portable infrared reflectometer capable of obtaining spectral reflectance of an opaque surface in the range of 2 to 25 microns using a Michelson-Type FTIR interferometer. This miniature interferometer collects many infrared spectra over a short period of time. It also allows the size of the instrument to be small such that spectra can be collected in the laboratory or in the field. Infrared spectra are averaged and integrated with respect to the room temperature black body spectrum to yield emittance at 300 K. Integrating with respect to other black body spectra yields emittance values at other temperatures. Absorption bands in the spectra may also be used for chemical species identification. The emittance of several samples was evaluated using this portable infrared reflectometer, an old infrared reflectometer equipped with dual rotating black body cavities, and a bench top thermal control vacuum chamber. Samples for evaluation were purposely selected such that a range of emittance values and thermal control material types would be represented, including polished aluminum, Kapton®, silvered Teflon®, and the inorganic paint Z-93-P. Results indicate an excellent linear relationship between the room temperature emittance calculated from infrared spectral data and the emittance obtained from the dual rotating black body cavities and thermal vacuum chamber. The prospect of using the infrared spectral date for chemical species identification will also be discussed.

de Groh, K. K., Jaworske, D. A., and Smith, D. C., "Optical Property Enhancement and Durability Evaluation of Heat Receiver Aperture Shield Materials", prepared for the 36th Aerospace Sciences Meeting and Exhibit sponsored by the American Institute of Aeronautics and Astronautics, Reno, Nevada, January 12-15, 1998.

Solar Dynamic (SD) power systems have been investigated by the National Aeronautics and Space Administration (NASA) for electrical power generation in space. As part of the International Space Station (ISS) program, NASA Glenn Research Center (GRC) teamed with the Russian Space Agency (RSA) to build a SD system to be flown on the Russian Space Station MIR. Under the US/Russian SD Flight Demonstration (SDFD) program, GRC worked with AlliedSignal Aerospace, the heat receiver contractor, on the development, characterization, and durability testing of materials to obtain appropriate optical and thermal properties for the SDFD heat receiver aperture shield. The aperture shield is composed of refractory metal multi-foil insulation (MFI) attached to an aperture back plate. Because of anticipated off-pointing periods, the aperture shield was designed to withstand the extreme temperatures that 80 W/cm² would produce. To minimize the temperature that the aperture shield will reach during off-pointing, it was desired for the aperture shield exterior layer to have a solar absorptance (a _s) to thermal emittance (e ) ratio as small as possible. In addition, a very low specular reflectance (r _s < 0.1) was also necessary, because reflected concentrated sunlight could cause overheating of the concentrator which is undesirable. Testing was conducted at GRC to evaluate pristine and optical property enhanced molybdenum and tungsten foils and screen covered foils. Molybdenum and tungsten foils were grit-blasted using silicon carbide or alumina grit under various grit-blasting conditions for optical property enhancement. Black rhenium coated tungsten foil was also evaluated. Tungsten, black rhenium-coated tungsten, and grit-blasted tungsten screens of various mesh sizes were placed over the pristine and grit-blasted foils for optical property characterization. Grit-blasting was found to be effective in decreasing the specular reflectance and absorptance/emittance ratio of the refractory foils. The placement of a screen further enhanced these optical properties, with a grit-blasted screen over a grit-blasted foil producing the best results. Based on the optical property enhancement results, samples were tested for atomic oxygen (AO) and vacuum heat treatment (VHT) durability. Grit-blasted (Al₂O₃ grit) 2 mil tungsten foil was chosen for the exterior layer of the SDFD heat receiver shield. A 0.007 in. diameter, 20x20 mesh tungsten screen was chosen to cover the tungsten foil. Based on these test results, a heat receiver aperture shield test unit has been built by Aerospace Design and Development (A.D.D.) with the screen covered grit-blast tungsten foil exterior layers. The aperture shield was tested in GRC's Solar Dynamic Ground Test Demonstration (SDGTD) system to verify the thermal and structural durability of the outer foil layers during an off-pointing period.

Jaworske, D. A., "Correlation of Predicted and Observed Optical Properties of Multi-Layer Thermal Control Coatings", Thin Solid Films 332, pp.30-33, 1998.

Thermal control coatings on spacecraft will be increasingly important as spacecraft grow smaller and more compact. New thermal control coatings will be needed to meet the demanding requirements of next generation spacecraft. Computer programs are now available to design optical coatings, and one such program was used to design several thermal control coatings consisting of alternating layers of WO₃ and SiO₂. The coatings were subsequently manufactured with electron beam evaporation and characterized with both optical and thermal techniques. Optical data were collected in both the visible region of the spectrum and the infrared. Solar absorptance values were predicted in the range of 0.177-0.196 and were observed in the range of 0.155-0.228. Infrared emittance values were predicted in the range of 0.074-0.083 and were observed optically in the range 0.048-0.093 and calorimetrically in the range of 0.069-0.100.

de Groh, K. K., Smith, D. C., Wheeler, D. R., and MacLachlan, B. J., "Effects of Ambient High Temperature Exposure on Alumina-Titania High Emittance Surfaces for Solar Dynamic Systems", prepared for the Space Technology and Applications International Forum, 1999, AIP CP 458, pp. 627-635, NASA TM-1998-208813.

Solar dynamic (SD) space power systems require durable, high emittance surfaces on a number of critical components, such as heat receiver interior surfaces and parasitic load radiator (PLR) elements. To enhance surface characteristics, an alumina-titania coating has been applied to 500 heat receiver thermal energy containment canisters and the PLR of NASA Glenn Research Center's (GRC) 2kW SD ground test demonstrator (GTD). The alumina-titania coating was chosen because it had been found to maintain its high emittance under vacuum at high temperatures for an extended period. However, preflight verification of SD system components, such as the PLR, require operation at ambient pressure and high temperatures. Therefore, the purpose of this research was to evaluate the durability of he alumina-titania coating at high temperature in air. Fifteen of sixteen alumina-titania coated Incoloy samples were exposed to high temperatures for various durations (2 to 32 hours). Samples were characterized prior to, and after, heat treatment for reflectance, solar absorptance, room temperature emittance, and emittance at 1200° F. Samples were also examined to detect physical defects and to determine surface chemistry using optical microscopy, scanning electron microscopy, operated with an energy dispersive spectroscopy (EDS) system, and x-ray photoelectron spectroscopy (XPS). Visual examination of the heat-treated samples showed a whitening of samples exposed to temperatures of 1000° F and above. Correspondingly, the optical properties of these samples degraded. A sample exposed to 1500° F for 24 hours had whitened and the thermal emittance at 1200° F had decreased from the non-heat treated value of 0.94 to 0.62. The coating on this sample had become embrittled, with spalling off the substrate noticeable at several locations. Based on this research it is recommended that preflight testing of SD components with alumina-titania coatings be restricted to temperatures no greater than 600° F in air to avoid optical degradation. Moreover, components with the alumina-titania coating are likely to experience optical property degradation with direct atomic oxygen exposure in space.

Jaworske, D. A., "Reflectivity of Silver and Silver-coated Substrates from 25°C to 800°C", prepared for the 32nd IECEC, Vol. 1, pp. 407-411, Honolulu, Hawaii, July, 1997.

A bench top facility was used to evaluate the reflectivity of several candidate coating-substrate combinations in vacuum at elevated temperatures. Silver was selected as the reflective coating of choice, while copper, nickel, electroless nickel on copper, and 304 stainless steel were selected as substrates. Pure silver, with no coating at all, was also evaluated. An optically flat silver-coated sapphire substrate was used as a standard. All metal substrates were either metallurgically polished or diamond turned to mirror finish prior to silver deposition. Silicon dioxide was used as a protective coating in most cases. Reflectivity measurements were made at room temperature in the visible range with a spectrophotometer, and at elevated temperatures up to 800°C with a helium-neon laser at 632 nm. Results from the high temperature reflectivity measurements will be presented.

Smith, C. A., Dever, J. A., and Jaworske, D. A., "Advances in Optical Property Measurements of Spacecraft Materials", presented at the 7th International Symposium on Materials in a Space Environment, Toulouse, France, June 16-20, 1997.

This report describes some of the instruments and experimental approaches available for measuring optical properties of thermal control materials. It also describes the instruments' uses in laboratory studies of the effects of combined contaminants and the space environment on these materials, and in the qualification of hardware for spacecraft. In recent years, several instruments for measurement of solar absorptance (a ) and infrared emittance (e ) have been introduced. These instruments offer improved speed, accuracy, and data-handling, all of which substantially improve the study of contaminated thermal control materials. A transient method for directly measuring material e is also described, and the results are compared with other instruments. In addition, our understanding of oxygen exposure effects on the ( of materials following contamination or exposure to simulated space conditions shows that oxygen exposure before measuring of e should be avoided.

The hemispherical total emissivity of two thermal control coatings, Z-93-P and black anodized aluminum, was calculated from hemispherical total reflectivity measured in wavelength range of 2 to 40 m m. These data were compared to hemispherical total emissivity values obtained on the same samples measured in a thermal vacuum chamber with a calorimetric technique. The comparison showed close agreement in the vicinity of room temperature and above, with differing trends at lower temperatures.

The hemispherical total emittances of several common thermal coatings tapes were evaluated calorimetrically over a wide temperature range. The calorimetric technique used here to evaluate thermal control materials allows a thermally isolated sample to cool solely through radiant heat transfer to a liquid nitrogen cooled cold wall. The mechanism of cooling is similar to that found in space, providing a functional evaluation of hemispherical total emittance. The five samples that were evaluated included several first and second surface mirrored Kapton tapes, one carbon-filled Kapton tape, and one black film. These tapes were adhesively bonded to an aluminum substrate, which provided the sensible heat for the calorimetric calculation. Temperature-time data were collected as heat flowed from the aluminum, through the adhesive layer, through the thermal control tape, and ultimately through the surface to the surroundings. Hemispherical total emittances were calculated over the temperature range of 100 to 400 K.

The emissivity of several different materials, including commercially available plastics, ceramics, and coatings, was evaluated by both optical and calorimetric means to select appropriate materials of construction for a microgravity combustion experiment. Samples of grit blasted black anodized aluminum, grit blasted black oxidized stainless steel, Bakelite^®, Noryl^®, Zelux^®, Macor^®, and Eccocoat^®, were all characterized in the range of 200 to 400 K. Several of these materials exceeded the combustion experiment emissivity requirement of 0.8 over a wide range of temperatures, suggesting their promising use as materials of construction. Emissivity values from all of the materials will be summarized, and those materials selected for the microgravity combustion experiment will be identified.

de Groh, K.K., Roig, D.M., Burke, C.A. and Shah D.R., "Performance and Durability of High Emittance Heat Receiver Surfaces for Solar Dynamic Power Systems", prepared for the 1994 ASME International Solar Energy Conference sponsored by the American Society of Mechanical Engineers, San Francisco, California, March 27-30, 1994. NASA/TM 106549

Haynes 188, a cobalt-based super-alloy, will be used to make thermal energy storage (TES) containment canisters for a 2 kW solar dynamic ground test demonstrator (SDGTD). Haynes 188 containment canisters with a high thermal emittance (e ) are desired for radiating heat away from local hot spots, improving the heat distribution, which will in turn improve canister service life. In addition to needing a high emittance, the surface needs to be durable in an elevated temperature, high vacuum (» 830° C, <10-7 torr) environment for an extended time period. Thirty-five Haynes 188 samples were exposed to 14 different types of surface modification techniques for emittance and vacuum heat treatment (VHT) durability enhancement. Optical properties were obtained for the modified surfaces. Emittance enhanced samples were exposed to VHT for up to 2692 hours at 827° C and <10-6 torr with integral thermal cycling. Optical properties were taken intermittently during exposure, and after final VHT exposure. The various surface modification treatments increased the emittance of pristine Haynes 188 from 0.11 to 0.86. Seven different surface modification techniques were found to provide surfaces which met the SDGTD receiver VHT durability requirement (e ³ 0.70 after 1000 hours). Of the 7 surface treatments, 2 were found to display excellent VHT durability: alumina-titania (AlTi) coatings (e = 0.85 after 2695 VHT hours) and zirconia-titania-yttria coatings (e = 0.86 after 2024.3 VHT hours). The AlTi coating was chosen for the e enhancement surface modification technique for the SDGTD receiver. Details of the alumina-titania coating and other Haynes 188 emittance surface modification techniques are discussed. Technology from this program will lead to successful demonstration of solar dynamic power for space applications, and has potential for applications in other systems requiring high emittance.

Thermal control surfaces are used in every spacecraft thermal management system to dissipate heat through heat transfer. This paper describes the thermal performance of several thermal control paints, coatings, and surfaces, as characterized by a calorimetric vacuum emissometer. The emissometer is designed to measure the functional emittance of a surface based on heat transfer from an underlying substrate to the surface and from the surface or near surface to a surrounding cold wall. Emittance measurements were made between 200 and 350 K. Polished aluminum, used here as a standard, was found to have a total hemispherical emittance of 0.06, as expected. A velvet black paint, also used as a standard, was found to have an emittance of 0.94 at room temperature. Other surfaces of interest included a polyurethane-based black paint designated Z-306, a highly polished 316L stainless steel, and an atomic beam-textured carbon-carbon composite.

A finite element analysis model of a transient technique used to measure the emittance of surface and coatings was developed and used to estimate the uncertainty in emittance. The dimensions used in the model matched the dimensions used in the design of a low temperature calorimetric vacuum emissometer being built to characterize the thermal properties of space power materials in the temperature range 173-673 K. Radiant energy from a quartz halogen lamp impinged on an aluminum sample that was coated with a thermal control coating and suspended in a liquid-nitrogen-cooled vacuum chamber by narrow gauge thermocouple wires. After removing the heat source, the temperature of the sample was monitored vs. time and the temperature-time curve was used to calculate the emittance. Factors contributing to the uncertainty in the emittance included uncertainties in time, temperature, area of the sample, heat capacity of the sample, and heat loss from the uncoated back of the sample. Heat losses from the thermocouple wires were found to be negligible. The total probable error in the emittance obtained from the low temperature calorimetric vacuum emissometer design was estimated to be less than 4% for emittance values greater than 0.5 at temperatures between 173 and 673 K.
 

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