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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 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS2005

Environmental Durability of Coated GRCop-84 Copper Alloys

An advanced Cu-8(at.%)Cr-4%Nb alloy developed at NASA's Glenn Research Center, and designated as GRCop-84, is currently being considered for use as liners in combustor chambers and nozzle ramps in NASA s future generations of reusable launch vehicles (RLVs). However, past experience has shown that unprotected copper alloys undergo an environmental attack called "blanching" in rocket engines using liquid hydrogen as fuel and liquid oxygen as the oxidizer. Potential for sulfidation attack of the liners in hydrocarbon-fueled engines is also of concern. As a result, protective overlay coatings alloys are being developed for GRCop-84. The oxidation behavior of several new coating alloys has been evaluated. GRCop-84 specimens were coated with several copper and nickel-based coatings, where the coatings were deposited by either vacuum plasma spraying or cold spraying techniques. Coated and uncoated specimens were thermally cycled in a furnace at different temperatures in order to evaluate the performance of the coatings. Additional studies were conducted in a high pressure burner rig using a hydrocarbon fuel and subjected to a high heat flux hydrogen-oxygen combustion flame in NASA s Quick Access Rocket Exhaust (QARE) rig. The performance of these coatings are discussed.

Raj, Sai V.↗
NASA NTRS2003

Delamination Mechanisms of Thermal and Environmental Barrier Coatings on SiC/SiC Ceramic Matrix Composites

Advanced ceramic thermal harrier coatings will play an increasingly important role In future gas turbine engines because of their ability to effectively protect the engine components and further raise engine temperatures. However, the coating durability issue remains a major concern with the ever-increasing temperature requirements. In this paper, thermal cyclic response and delamination failure modes of a ZrO2-8wt%Y2O3 and mullite/BSAS thermaVenvironmenta1 barrier coating system on SiC/SiC ceramic matrix composites were investigated using a laser high-heat-flux technique. The coating degradation and delamination processes were monitored in real time by measuring coating apparent conductivity changes during the cyclic tests under realistic engine temperature and stress gradients, utilizing the fact that delamination cracking causes an apparent decrease in the measured thermal conductivity. The ceramic coating crack initiation and propagation driving forces under the cyclic thermal loads, in conjunction with the mechanical testing results, will be discussed.

Zhu, Dongming↗
NASA NTRS1990

Delamination Mechanisms of Thermal and Environmental Barrier Coatings on SiC/SiC Ceramic Matrix Composites

Advanced ceramic thermal barrier coatings will play an increasingly important role in future gas turbine engines because of their ability to effectively protect the engine components and further raise engine temperatures. However, the coating durability issue remains a major concern with the ever-increasing temperature requirements. In this paper, thermal cyclic response and delamination failure modes of a ZrO2-8wt%Y2O3 and mullite/BSAS thermal/environmental barrier coating system on SiC/SiC ceramic matrix composites were investigated using a laser high-heat-flux technique. The coating degradation and delamination processes were monitored in real time by measuring coating apparent conductivity changes during the cyclic tests under realistic engine temperature and stress gradients, utilizing the fact that delamination cracking causes an apparent decrease in the measured thermal conductivity. The ceramic coating crack initiation and propagation driving forces under the cyclic thermal loads, in conjunction with the mechanical testing results, will be discussed.

Zhu, Dongming↗
NASA NTRS1990

Surface Cracking and Interface Reaction Associated Delamination Failure of Thermal and Environmental Barrier

In this paper, surface cracking and interface reactions of a ZrO2-8wt%Y2O3 and mullite/BSAS/Si thermal and environmental barrier coating system on SiC/SiC ceramic matrix composites were characterized after long-term combined laser thermal gradient and furnace cyclic tests in a water vapor containing environment. The surface cracking was analyzed based on the coating thermal gradient sintering behavior and thermal expansion mismatch stress characteristics under the thermal cyclic conditions. The interface reactions that were largely enhanced by the coating surface cracking in the water vapor environment were investigated in detail, and the reaction phases were identified for the coating system after the long-term exposure. The accelerated coating delamination failure was attributed to the increased delamination driving force under the thermal gradient cyclic loading and the reduced interface adhesion due to the detrimental interface reactions. The coating design issues will also be discussed based on the observed failure mechanisms under the high-heat-flux test conditions.

Zhu, Dongming↗
NASA NTRS2001

Thermal Barrier Coatings for Airfoil Applications

Thermal barrier coatings having sigruficantly reduced thermal conductivities are being developed under this program. The conductivities, after 20 hours of sintering under high heat flux exposure in a unique laser rig test, are reduced to nearly 1/3 of the state-of-the-art baseline value for the plasma sprayed coatings and to 1/2 of the baseline for the physical vapor-deposited coatings. The coatings consist of zirconia-yttria with relatively small additions of certain rare earth oxides. Further conductivity reductions and cyclic life optimization are planned for the coming year.

Miller, Robert A.↗
NASA NTRS2003

Low Thermal Conductivity Thermal Barrier Coatings Developed

Thermal barrier coatings (TBCs) are used extensively in modern gas turbine engines to thermally insulate air-cooled metallic components from the hot gases in the engine. These coatings typically consist of a zirconia-yttria ceramic that has been applied by either plasma spraying or physical vapor deposition. Future engines will rely even more heavily on TBCs and will require materials that have even higher temperature capability with improved insulation (i.e., lower thermal conductivity even after many hours at high temperature). This report discusses new TBCs that have been developed with these future requirements in mind. The Ultra-Efficient Engine Technology Program at the NASA Glenn Research Center is funding this effort, which has been conducted primarily at Glenn with contractor support (GE and Howmet) for physical vapor deposition. As stated, the new TBC not only had to be more insulating but the insulation had to persist even after many hours of exposure-that is, the new TBC had to have both lower conductivity and improved sintering resistance. A new type of test rig was developed for this task. This new test approach used a laser to deliver a known high heat flux in an essentially uniform pattern to the surface of the coating, thereby establishing a realistic thermal gradient across its thickness. This gradient was determined from surface and backside pyrometry; and since the heat flux and coating thickness are known, this permitted continuous monitoring of thermal conductivity. Thus, this laser rig allowed very efficient screening of candidate low-conductivity, sinter-resistant TBCs. The coating-design approach selected for these new low-conductivity TBCs was to identify oxide dopants that had the potential to promote the formation of relatively large and stable groupings of defects known as defect clusters. This approach was used because it was felt that such clusters would reduce conductivity while enhancing stability. The approach proved to be successful: low-conductivity TBCs having improved sintering resistance were developed.

Zhu, Dongming↗
NASA NTRS2003

Glenn-HT Code Validated for Complex Turbine Blade Cooling Passage

This work is motivated by the need to accurately predict heat transfer in turbomachinery. For efficient gas turbine operation, flow temperatures in the hot gas path exceed acceptable metal temperatures in many regions of the engine. So that the integrity of the parts can be maintained for an acceptable engine life, the parts must be cooled. Efficient cooling schemes require accurate heat transfer prediction to minimize regions that are overcooled and, even more importantly, to ensure adequate cooling in high-heat-flux regions.

Rigby, David L.↗
NASA NTRS2005

Low Conductivity Thermal Barrier Coatings

Thermal barrier coatings will be more aggressively designed to protect gas turbine engine hot-section components in order to meet future engine higher fuel efficiency and lower emission goals. In this presentation, thermal barrier coating development considerations and requirements will be discussed. An experimental approach is established to monitor in real time the thermal conductivity of the coating systems subjected to high-heat-flux, steady-state and cyclic temperature gradients. Advanced low conductivity thermal barrier coatings have also been developed using a multi-component defect clustering approach, and shown to have improved thermal stability. The durability and erosion resistance of low conductivity thermal barrier coatings have been improved utilizing advanced coating architecture design, composition optimization, in conjunction with more sophisticated modeling and design tools.

Zhu, Dong-Ming↗
NASA NTRS2006

Comparison of AMZIRC and GRCop-84

The mechanical properties of two copper alloys with high thermal conductivity, GRCop-84 and AMZIRC, were compared. These are competing alloys in high temperature, high heat flux applications such as rocket nozzles. The GRCop-84 data presented was taken from previous work. The results of new tensile, creep, and compression tests on AMZIRC are presented. Tests were done on as-received hard drawn material, and on material that had been subjected to a heat treatment designed to simulate a brazing operation at 935 C. As-received AMZIRC was found to have excellent properties at temperatures below 550 C, with room temperature yield and ultimate tensile strengths of about 500 MPa, and ductile failures. By comparison, GRCop-84 s room temperature tensile yield and ultimate strengths are about 200 and 380 MPa respectively. However, the simulated brazing heat treatment substantially decreased the mechanical properties of AMZIRC; and the strength of as-received AMZIRC dropped precipitously as test temperatures exceeded 500 C. The properties of GRCop-84 were not significantly affected by the 935 C heat treatment. As a result, there appear to be advantages to GRCop- 84 over AMZIRC if use or processing temperatures of greater than 500 C are expected. Tensile creep tests were done at 500 and 650 C. At these temperatures, the creep properties of GRCop-84 were superior to AMZIRC s. At equivalent rupture life and stress, GRCop-84 was found to have a 150 C temperature advantage over AMZIRC; for equivalent rupture life and temperature GRCop-84 was two times stronger.

de Groh, Henry C. III↗
NASA NTRS2006

Lightweight Ablative and Ceramic Thermal Protection System Materials for NASA Exploration Systems Vehicles

As a collaborative effort among NASA Centers, the "Lightweight Nonmetallic Thermal Protection Materials Technology" Project was set up to assist mission/vehicle design trade studies, to support risk reduction in thermal protection system (TPS) material selections, to facilitate vehicle mass optimization, and to aid development of human-rated TPS qualification and certification plans. Missions performing aerocapture, aerobraking, or direct aeroentry rely on advanced heatshields that allow reductions in spacecraft mass by minimizing propellant requirements. Information will be presented on candidate materials for such reentry approaches and on screening tests conducted (material property and space environmental effects tests) to evaluate viable candidates. Seventeen materials, in three classes (ablatives, tiles, and ceramic matrix composites), were studied. In additional to physical, mechanical, and thermal property tests, high heat flux laser tests and simulated-reentry oxidation tests were performed. Space environmental effects testing, which included exposures to electrons, atomic oxygen, and hypervelocity impacts, was also conducted.

Valentine, Peter G.↗
NASA NTRS2007

Development of GRCop-84 for Rocket Engine Applications

GRCop-84 (Cu-8 at.% Cr-4 at.% Nb) has been under development at the NASA Glenn Research Center for several years. The alloy possesses a unique combination of good thermal conductivity, high elevated temperature strength, long creep lives and long low cycle fatigue lives. The alloy is also more oxidation resistant than pure copper and most competitive alloys. The combination of properties has attracted attention from major rocket engine manufacturers who are interested in the alloy for the combustion chamber liner in their next generation of regeneratively cooled engines. It is also a strong candidate for various components in hypersonic vehicles, which also experience very high heat flux conditions. This presentation will discuss the alloy design strategies used to develop GRCop-84 and the various processing routes available for manufacturing components. The microstructure and mechanical properties of the alloy will be reviewed, as well as the results of actual hot fire testing of subscale rocket combustion chambers. The use of environmental and thermal barrier coatings to extend the performance to even higher levels will also be discussed.

Ellis, David L.↗
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 NTRS2008

Thermal and Environmental Barrier Coating Development for Advanced Propulsion Engine Systems

Ceramic thermal and environmental barrier coatings (TEBCs) are used in gas turbine engines to protect engine hot-section components in the harsh combustion environments, and extend component lifetimes. Advanced TEBCs that have significantly lower thermal conductivity, better thermal stability and higher toughness than current coatings will be beneficial for future low emission and high performance propulsion engine systems. In this paper, ceramic coating design and testing considerations will be described for turbine engine high temperature and high-heat-flux applications. Thermal barrier coatings for metallic turbine airfoils and thermal/environmental barrier coatings for SiC/SiC ceramic matrix composite (CMC) components for future supersonic aircraft propulsion engines will be emphasized. Further coating capability and durability improvements for the engine hot-section component applications can be expected by utilizing advanced modeling and design tools.

Zhu, Dongming↗
NASA NTRS2007

Testing Lunar Return Thermal Protection Systems using Sub-Scale Flight Test Vehicles

A key objective of NASA's Vision for Space Exploration is to revisit the lunar surface. Such an ambitious goal requires the development of a new human-rated spacecraft, the Orion Crew Exploration Vehicle (CEV), to ferry crews to low earth orbit and to the moon. The successful conclusion of both types of missions will require a thermal protection system (TPS) capable of protecting the vehicle and crew from the extreme heat of atmospheric reentry. As a part of the TPS development, various materials are being tested in arcjet tunnels; however, the combined lunar return aerothermal environment of high heat flux, shear stress, and surface pressure cannot be duplicated using only existing ground test facilities. To ensure full TPS qualification, a flight test program using sub-scale Orion capsules has been proposed to test TPS materials and heat shield construction techniques under the most stressing combination of lunar return aerothermal environments. Originally called Testing Of Reentry Capsule Heat Shield, or TORCH, but later renamed LEX, for Lunar Reentry Experiment, the proposed flight test program is presented along with the driving requirements and descriptions of the vehicle and the TPS instrumentation suite slated to conduct in-flight measurements.

heatshield↗
NASA NTRS2006

Tantalum-Based Ceramics for Refractory Composites

A family of tantalum-based ceramics has been invented as ingredients of high-temperature composite insulating tiles. These materials are suitable for coating and/or permeating the outer layers of rigid porous (foam-like or fibrous) ceramic substrates to (1) render the resulting composite ceramic tiles impervious to hot gases and (2) enable the tiles to survive high heat fluxes at temperatures that can exceed 3,000 F ( 1,600 C).

Stewart, David A.↗
NASA NTRS2011

Mechanical Properties and Durability of Advanced Environmental Barrier Coatings in Calcium-Magnesium-Alumino-Silicate Environments

Environmental barrier coatings are being developed and tested for use with SiC/SiC ceramic matrix composite (CMC) gas turbine engine components. Several oxide and silicate based compositons are being studied for use as top-coat and intermediate layers in a three or more layer environmental barrier coating system. Specifically, the room temperature Vickers-indentation-fracture-toughness testing and high-temperature stability reaction studies with Calcium Magnesium Alumino-Silicate (CMAS or "sand") are being conducted using advanced testing techniques such as high pressure burner rig tests as well as high heat flux laser tests.

Miladinovich, Daniel S.↗
NASA NTRS2011

Development of Low Density Flexible Carbon Phenolic Ablators

Phenolic Impregnated Carbon Ablator (PICA) was the enabling TPS material for the Stardust mission where it was used as a single piece heatshield. PICA has the advantages of low density (0.27g/cm3) coupled with efficient ablative capability at high heat fluxes. Under the Orion program, PICA was also shown to be capable of both ISS and lunar return missions however some unresolved issues remain for its application in a tiled configuration for the Orion-specific design. In particular, the problem of developing an appropriate gap filler resulted in the Orion program selecting AVCOAT as the primary heatshield material over PICA. We are currently looking at alternative architectures to yield flexible and more conformal carbon phenolic materials with comparable densities to PICA that will address some of the design issues faced in the application of a tiled PICA heat shield. These new materials are viable TPS candidates for upcoming NASA missions and as material candidates for private sector Commercial Orbital Transportation Services (COTS). This presentation will discuss flexible alternatives to PICA and include preliminary mechanical and thermal properties as well as arc jet and LHMEL screening test results.

Stackpole, Mairead↗
NASA NTRS2012

Transpiration-Cooled Spacecraft-Insulation-Repair Fasteners

Transpiration-cooled fasteners are proposed that operate like an open-loop heat pipe (self-tapping screws, bolts, and spikes) for use in on-orbit repair of thermal- insulation of a space shuttle or other spacecraft. By limiting the temperature rise of such a fastener and of the adjacent repair material and thermal protection system, the transpiration cooling would contribute to the ability of the repair to retain its strength and integrity in the high-heat-flux, oxidizing environment of reentry into the atmosphere of the Earth. A typical fastener according to the proposal would include a hollow refractory-metal, refractory-composite, or ceramic screw or bolt, the central cavity of which would be occupied by a porous refractory- metal or ceramic plug that would act as both a reservoir and a wick for a transpirant liquid. The plug dimensions, the plug material, and the sizes of the pores would be chosen in conjunction with the transpirant liquid so that (1) capillary pumping could be relied upon to transport the liquid to the heated surface, where the liquid would be vaporized, and (2) the amount of liquid would suffice for protecting against the anticipated heat flux and integrated heat load.

Camarda, Charles J.↗