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Solar thermal conversion performance is assessed in this paper for representative point focusing distributed systems. Trade-off comparisons are made in terms of concentrator quality, solar receiver operating temperature, and power conversion efficiency. Normalized system performance is presented on a unit concentrator area basis for integrated annual electric energy production.
Gas filled insulation panels (GFP) are very light weight and compact (when uninflated) advanced insulation products. GFPs consist of multiple layers of thin, low emittance (low-e) metalized aluminum. When expanded, the internal, low-e aluminum layers form a honeycomb structure. These baffled polymer chambers are enveloped by a sealed barrier and filled with either air or a low-conductivity gas. The sealed exterior aluminum foil barrier films provide thermal resistance, flammability protection, and properties to contain air or a low conductivity inert gas. This product was initially developed with a grant from the U.S. Department of Energy. The unexpanded product is nearly flat for easy storage and transport. Therefore, transportation volume and weight of the GFP to fill unit volume of wall cavity is much smaller compared to that of other conventional insulation products. This feature makes this product appealing to use at Army Contingency Basing, when transportation cost is significant compared to the cost of materials. The objective of this study is to evaluate thermal performance of walls, similar to those used at typical Barracks Hut (B-Hut) hard shelters, when GFPs are used in the wall cavities. Oak Ridge National Laboratory (ORNL) tested performance of the wall in the rotatable guarded hotbox (RGHB) according to the ASTM C 1363 standard test method.
The procedures used and the results obtained during the evaluation test program on the S. N. 1, (liquid) solar collector are presented. The flat plate collector uses water as the working fluid. The absorber plate is aluminum with plastic tubes coated with urethane black. The glazing consists of .040 in fiberglass reinforced polyester. The collector weight is 78.5 pounds with overall external dimensions of approximately 50.3in. x 98.3in. x 3.8in. The following information is given: thermal performance data under simulated conditions, structural behavior under static loading, and the effects of long term exposure to natural weathering. These tests were conducted using the MSFC Solar Simulator.
Maritime shipping of vast quantities of liquid hydrogen (LH 2 ) will be necessary to facilitate a global hydrogen ecosystem; with some studies estimating volumes up to 172,000 m 3 for individual tanker ships, and requiring stationary storage tanks at terminals of 50,000 m 3 to 100,000 m 3 —ten to fifteen times larger than the current largest tank, located at launch pad B at NASA Kennedy Space Center (KSC). Such a radical scale-up will push the boundary of traditional, vacuum-insulated tank designs. Hence, a potential need exists for non-vacuum solutions, which necessitates exploring the thermal performance of insulation materials in non-condensable background gasses at LH 2 temperatures, namely helium and hydrogen. Testing of two bulk-fill insulation materials common to large LH 2 storage tanks, perlite and glass bubbles, in helium and hydrogen was recently conducted by the Cryogenics Test Laboratory at KSC using the Cryostat-100 liquid nitrogen boiloff calorimeter per the ASTM C1774 standard methodology. Effective thermal conductivity (k e ) and heat flux (q) results for each insulation/gas combination are presented across the full vacuum range, as well as thermal profiles through the insulation thickness between 78 K and 293 K.
Procedures used and results obtained during the evaluation test program on a flat plate collector which uses water as the working fluid are discussed. The absorber plate is copper tube soldered to copper fin coated with flat black paint. The glazing consists of two plates of Lo-Iron glass; the insulation is polyurethane foam. The collector weight is 242.5 pounds with overall external dimensions of approximately 48.8 in. x 120.8 in. x 4.1 in. The test program was conducted to obtain thermal performance data before and after 34 days of weather exposure test.
NASA launched two rovers in June and July of 2003 as a part of the Mars Exploration Rover (MER) project. MER-A (Spirit) landed on Mars in Gusev Crater at 15 degrees South latitude and 175 degree East longitude on January 4, 2004 (Squyres, et al., Dec. 2004)). MER-B (Opportunity) landed on Mars in Terra Meridiani at 2 degrees South latitude and 354 degrees East longitude on January 25, 2004 (Squyres, et al., August 2004) Both rovers have well exceeded their design lifetime (90 Sols) by more than a factor of 4. Spirit and Opportunity are still healthy and continue to execute their roving science missions at the time of this writing. This paper discusses rover flight thermal performance during the surface missions of both vehicles, covering roughly the time from the MER-A landing in late Southern Summer (Ls = 328, Sol 1A) through the Southern Winter solstice (Ls = 90, Sol 255A) to nearly Southern Vernal equinox (Ls = 160 , Sol 398A).
The Geoscience Laser Altimeter System (GLAS) instrument is NASA Goddard Space Flight Center's first application of Loop Heat Pipe technology that provides selectable/stable temperature levels for the lasers and other electronics over a widely varying mission environment. GLAS was successfully launched as the sole science instrument aboard the Ice, Clouds, and Land Elevation Satellite (ICESat) from Vandenberg AFB at 4:45pm PST on January 12, 2003. After SC commissioning, the LHPs started easily and have provided selectable and stable temperatures for the lasers and other electronics. This paper discusses the thermal development background and testing, along with details of early flight thermal performance data.
The procedures used and the results obtained during the evaluation test program on the Solar Unlimited, Inc., Suncatcher SH-11 (liquid) solar collector are presented. The flat-plate collector case assembly is made of .08 inch aluminum 3003 H14 riveted with fiberglass board insulation. The absorber consists of collared aluminum fins mechanically bonded to 3/8 inch copper tubing and coated with 3M Nextel black. Water is used as the working fluid. The glazing is made of a single glass, 1/8 inch water white, tempered and antireflective. The collector weight is 85 pounds with overall external dimensions of about 35.4 in x 82.0 in x 4.0 in. Thermal performance data on the Solar Unlimited Suncatcher SH-11 solar collector under simulated conditions were conducted using the MSFC Solar Simulator.
A new geometry of the dew-point indirect evaporative cooler structure was proposed which enables uniform distribution of water in the working channels and application in traditional air handling units. This arrangement requires a complicated structure of the product channel in which the air is cooled. Further, the airflow direction changes and complicated structure it is associated with local losses affecting the total pressure loss. High-pressure drop limits the use of this dew point evaporative cooler as an alternative source of cooling. Therefore, a verified method based on numerical fluid dynamics (CFD) was used to determine the pressure drop and air distribution in the exchanger. The combination of the CFD method and modified epsilon-NTU allows the thermal performance of the device including the uniformity of air distribution, cooling capacity, and coefficient of performance to be determined. Finally, the appropriate exchanger dimensions, i.e., the channel height, or final distribution regulation may be found depending on the exchanger application.
The NASA Cryogenics Test Laboratory at Kennedy Space Center conducted long-term testing of SOFI materials under actual-use cryogenic conditions with Cryostat-4. The materials included in the testing were NCFI 24-124 (acreage foam), BX-265 (close-out foam, including intertank flange and bipod areas), and a potential alternate material, NCFI 27-68, (acreage foam with the flame retardant removed). Specimens of these materials were placed at two locations: a site that simulated aging (the Vehicle Assembly Building [VAB]) and a site that simulated weathering (the Atmospheric Exposure Test Site [beach site]). After aging/weathering intervals of 3, 6, and 12 months, the samples were retrieved and tested for their thermal performance under cryogenic vacuum conditions with test apparatus Cryostat-4.
Additive manufacturing has transformed thermal management by enabling the production of complex, optimized geometries that conventional manufacturing methods cannot achieve. This study investigates the single-phase convective heat transfer performance of gyroid triply periodic minimal surface (TPMS) lattice structures with functional porosity. TPMS structures provide high surface area to volume ratios and are amenable to 3D printing. A gyroid numerical model was created and validated against an existing experimental study with a similar feature size to the investigated geometries. The TPMS structure has a periodic width of 1.6 mm, a length of 10 mm, and a height of 4 mm, with a functional porosity ranging from 0.5 to 0.8, decreasing with distance from the heated surface. Three different flow configurations were examined for an inlet fluid temperature of 70 °C. The inlet velocities range from 0.01 to 1.2 m/s, corresponding to a Reynolds number range of 10–900 with a heat flux of 50 W/cm 2 applied at the base. AmpCool ® AC-110 dielectric fluid (Prandtl number 59.5) was used as the coolant. Thermal performance and friction characteristics were studied for the three flow orientations. The parallel flow configuration was identified as the most efficient for heat removal. A detailed analysis of the numerical results highlights the underlying physics behind the thermal performance differences among the flow configurations.
The Terra spacecraft is the flagship of NASA's Earth Science Enterprise. It provides global data on the state of atmosphere, land and oceans, as well as their interactions with solar radiation and one another. Three Terra instruments utilize Capillary Pumped Heat Transport System (CPHTS) for temperature control: Each CPHTS, consisting of two capillary pumped loops (CPLs) and several heat pipes and electrical heaters, is designed for instrument heat loads ranging from 25W to 264W. The working fluid is ammonia. Since the launch of the Terra spacecraft, each CPHTS has been providing a stable interface temperature specified by the instrument under all modes of spacecraft and instrument operations. The ability to change the CPHTS operating temperature upon demand while in service has also extended the useful life of one instrument. This paper describes the design and on-orbit performance of the CPHTS thermal systems.
In thermal energy storage systems, phase change materials (PCMs) are widely used to provide efficient storing and extraction of thermal energy. However, most PCMs have low thermal conductivities, which limit the heat transfer of the latent energy inward/outward. In this study, the impacts of various fiber inserts on the solidification and melting processes with PCM were investigated by experiments and simulations. It was found that long anisotropic carbon fibers with high thermal conductivities can not only increase the heat transfer within PCM by enlarging hot-cold interface areas, but also act as crystallization nuclei and increase the solidification of PCM during the charging process. Recalescence of PCM during solidification was observed. COMSOL models were used to understand the thermal transfer mechanisms. COMSOL Multiphysics models of a single carbon fiber filament in PCM showed that the anisotropy in thermal conductivities controlled the radial and axial heat transfer ratios along its fiber axis and can enhance both the solidification and melting processes.
The challenging range of proposed landing sites for the Mars Science Laboratory Rover requires a rover thermal management system that is capable of keeping temperatures controlled across a wide variety of environmental conditions. On the Martian surface where temperatures can be as cold as -123 degrees Centigrade and as warm as 38 degrees Centigrade, the Rover relies upon a Mechanically Pumped Fluid Loop (MPFL) and external radiators to maintain the temperature of sensitive electronics and science instruments within a -40 degrees Centigrade to 50 degrees Centigrade range. The MPFL also manages significant waste heat generated from the Rover power source, known as the Multi Mission Radioisotope Thermoelectric Generator (MMRTG). The MMRTG produces 110 Watts of electrical power while generating waste heat equivalent to approximately 2000 Watts. Two similar Heat Exchanger (HX) assemblies were designed to both acquire the heat from the MMRTG and radiate waste heat from the onboard electronics to the surrounding Martian environment. Heat acquisition is accomplished on the interior surface of each HX while heat rejection is accomplished on the exterior surface of each HX. Since these two surfaces need to be at very different temperatures in order for the MPFL to perform efficiently, they need to be thermally isolated from one another. The HXs were therefore designed for high in-plane thermal conductivity and extremely low through-thickness thermal conductivity by using aerogel as an insulator inside composite honeycomb sandwich panels. A complex assembly of hand welded and uniquely bent aluminum tubes are bonded onto the HX panels and were specifically designed to be easily mated and demated to the rest of the Rover Heat Recovery and Rejection System (RHRS) in order to ease the integration effort. During the cruise phase to Mars, the HX assemblies serve the additional function of transferring heat from the Rover MPFL to the separate Cruise Stage MPFL so that heat generated deep inside the Rover can be dissipated via the Cruise Stage radiators. Significant fabrication challenges had to be overcome in order to make the HX design a reality. The cruise phase thermal performance of the Rover HXs was verified in the cruise phase system level thermal vacuum test that was performed at JPL in January of 2009. The Rover HXs were modeled in I-DEAS TMG and predictions are compared to actual data from the test.
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A large telescope aperture, stringent thermal stability and temperature range requirements, and a passively-cooled 1500K module presented major challenges in thermal design and hardware fabrication of this Small Explorer satellite. This paper reviews briefly the thermal design of the SWAS science instrument, and examines the first three months of on-orbit thermal history. Measured temperatures for both the science payload and the spacecraft module and solar arrays are compared with those predicted by the correlated analytical model. Similarities and differences are interpreted in terms of the major uncertainties remaining after thermal-balance testing, especially those of MLI performance and telescope aperture properties. Review of the thermal model adequacy and thermal design verification are included to suggest improvements in the thermal design process for future missions.
A large telescope aperture, stringent thermal stability and temperature range requirements, and a passively-cooled 150 K module presented major challenges in thermal design and hardware fabrication of this Small Explorer satellite. This paper reviews briefly the thermal design of the SWAS science instrument, and examines the first three months of on-orbit thermal history. Measured temperatures for both the science payload and the spacecraft module and solar arrays are compared with those predicted by the correlated analytical model. Similarities and differences are interpreted in terms of the major uncertainties remaining after thermal-balance testing, especially those of MLI performance and telescope aperture properties. Review of the thermal model adequacy and thermal design verification are included to suggest improvements in the thermal design process for future missions.
Here, this paper presents experimental results from conduction-cooled thermal testing of a ReBCO canted cosine theta (CCT) magnet (C2), originally designed and fabricated at LBNL using CORC cables. While the performance of the coil under liquid helium and nitrogen environments has been previously established, this study explores its behavior under conduction cooling using a large test cryostat at The Ohio State University. The magnet, measuring 613 mm in length and weighing ∼75 kg, consists of four helical layers wound with ReBCO-based CORC wire and was thermally anchored to a copper cold ring supported by a G-10 strongback. Cooling was provided by two Sumitomo RDK-415D cryocoolers, offering a combined 3 W at 4.2 K and 150 W at 77 K. Multiple thermal cycles were performed, with cooldown durations of up to 45 hours. Final base temperatures of approximately 10.8 K (at the coil edge) and 12.0 K (at the coil center) were achieved, with an axial temperature difference of approximately 1.2 K. The warm-up period extended over approximately 24.6 hours. Voltage measurements from all four layers were recorded during cooldown and warmup. The system demonstrated stable cooldown performance, repeatable gradients, and good thermal anchoring. These results support the feasibility of conduction cooling in large-scale HTS magnets, aligning with broader goals for “green” cryogen-free accelerator technologies and paving the way for more sustainable, scalable, and energy-efficient high-field magnet systems in next-generation particle accelerators.