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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 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

Oxidation of Carbon Fibers in a Cracked Ceramic Matrix Composite Modeled as a Function of Temperature

The oxidation model simulates the oxidation of the reinforcing carbon fibers within a ceramic matrix composite material containing as-fabricated microcracks. The physics-based oxidation model uses theoretically and experimentally determined variables as input for the model. The model simulates the ingress of oxygen through microcracks into a two-dimensional plane within the composite material. Model input includes temperature, oxygen concentration, the reaction rate constant, the diffusion coefficient, and the crack opening width as a function of the mechanical and thermal loads. The model is run in an iterative process for a two-dimensional grid system in which oxygen diffuses through the porous and cracked regions of the material and reacts with carbon in short time steps. The model allows the local oxygen concentrations and carbon volumes from the edge to the interior of the composite to be determined over time. Oxidation damage predicted by the model was compared with that observed from microstructural analysis of experimentally tested composite material to validate the model for two temperatures of interest. When the model is run for low-temperature conditions, the kinetics are reaction controlled. Carbon and oxygen reactions occur relatively slowly. Therefore, oxygen can bypass the carbon near the outer edge and diffuse into the interior so that it saturates the entire composite at relatively high concentrations. The kinetics are limited by the reaction rate between carbon and oxygen. This results in an interior that has high local concentrations of oxygen and a similar amount of consumed carbon throughout the cross section. When the model is run for high-temperature conditions, the kinetics are diffusion controlled. Carbon and oxygen reactions occur very quickly. The carbon consumes oxygen as soon as it is supplied. The kinetics are limited by the relatively slow rate at which oxygen is supplied in comparison to the relatively fast rate at which carbon and oxygen reactions occur. This results in a sharp gradient in oxygen concentration from the edge where it is supplied to the nearest source of carbon, which is where the oxygen is quickly consumed. A moving reaction front is seen in which the outlaying carbon is consumed before the next inner layer of carbon begins to react.

Halbig, Michael C.↗
NASA NTRS1989

Two-layer thermal barrier coating. I - Effects of composition and temperature on oxidation behavior and failure

The effects of 21 bond and 2 ceramic coating compositions on the specific mass gain, internal crack location at failure, and life of a two-layer thermal barrier coating (TBC) were studied by cyclic testing in a furnace. MAR-M 200 + Hf alloy specimens were completely coated with bond and thermal barrier (ceramic) coatings. Both coatings were applied by air plasma spray deposition. Cyclic test data were obtained at 1110, 1160, and 1220 deg C. The data show that the specific mass gain and the TBC life are significantly affected by the composition of the bond coating and the temperature and only slightly affected by the composition of the ceramic coating.

Stecura, Stephan↗
NASA NTRS2010

Nondestructive Evaluation (NDE) for Characterizing Oxidation Damage in Cracked Reinforced Carbon-Carbon

In this study, coated reinforced carbon-carbon (RCC) samples of similar structure and composition as that from the NASA space shuttle orbiter's thermal protection system were fabricated with slots in their coating simulating craze cracks. These specimens were used to study oxidation damage detection and characterization using nondestructive evaluation (NDE) methods. These specimens were heat treated in air at 1143 C and 1200 C to create cavities in the carbon substrate underneath the coating as oxygen reacted with the carbon and resulted in its consumption. The cavities varied in diameter from approximately 1 to 3mm. Single-sided NDE methods were used because they might be practical for on-wing inspection, while X-ray micro-computed tomography (CT) was used to measure cavity sizes in order to validate oxidation models under development for carbon-carbon materials. An RCC sample having a naturally cracked coating and subsequent oxidation damage was also studied with X-ray micro-CT. This effort is a follow-on study to one that characterized NDE methods for assessing oxidation damage in an RCC sample with drilled holes in the coating.

Roth, Don J.↗
NASA NTRS1995

Microcracking in composite laminates under thermal and mechanical loading

Composites used in space structures are exposed to both extremes in temperature and applied mechanical loads. Cracks in the matrix form, changing the laminate thermoelastic properties. The goal of the present investigation is to develop a predictive methodology to quantify microcracking in general composite laminates under both thermal and mechanical loading. This objective is successfully met through a combination of analytical modeling and experimental investigation. In the analysis, the stress and displacement distributions in the vicinity of a crack are determined using a shear lag model. These are incorporated into an energy based cracking criterion to determine the favorability of crack formation. A progressive damage algorithm allows the inclusion of material softening effects and temperature-dependent material properties. The analysis is implemented by a computer code which gives predicted crack density and degraded laminate properties as functions of any thermomechanical load history. Extensive experimentation provides verification of the analysis. AS4/3501-6 graphite/epoxy laminates are manufactured with three different layups to investigate ply thickness and orientation effects. Thermal specimens are cooled to progressively lower temperatures down to -184 C. After conditioning the specimens to each temperature, cracks are counted on their edges using optical microscopy and in their interiors by sanding to incremental depths. Tensile coupons are loaded monotonically to progressively higher loads until failure. Cracks are counted on the coupon edges after each loading. A data fit to all available results provides input parameters for the analysis and shows them to be material properties, independent of geometry and loading. Correlation between experiment and analysis is generally very good under both thermal and mechanical loading, showing the methodology to be a powerful, unified tool. Delayed crack initiation observed in a few cases is attributed to a lack of preexisting flaws assumed by the analysis. Some interactions between adjacent ply groups are attributed to local stress concentrations. These two effects are not captured by the analysis due to its global nature. The analysis is conservative in these cases and agrees well with data after the observed onset of cracking.

Maddocks, Jason R.↗
NASA NTRS1998

Degradation of Teflon(tm) FEP Following Charged Particle Radiation and Rapid Thermal Cycling

During the Second Servicing Mission (SM2) of the Hubble Space Telescope (HST) severe degradation was observed on the outer layer of the thermal control blankets. Astronaut observations and photographs revealed large cracks in the metallized Teflon FEP (fluorinated ethylene propylene), the outer layer of the multi-layer insulation (MLI), in many locations around the telescope. In an effort to understand what elements of the space environment might cause such damage, pristine Teflon' FEP was tested for durability to radiation and thermal cycling. Specimens were subjected to electron and proton fluences comparable to those experienced by HST and were subsequently thermal cycled in a custom-built rapid thermal cycle chamber. Tensile tests of the specimens showed that radiation followed by thermal cycling significantly reduced the ultimate strength and elongation of Teflon FEP.

Townsend, Jacqueline A.↗
NASA NTRS2009

Nondestructive Evaluation (NDE) for Characterizing Oxidation Damage in Cracked Reinforced Carbon-Carbon (RCC)

In this study, coated reinforced carbon-carbon (RCC) samples of similar structure and composition as that from the NASA space shuttle orbiter's thermal protection system were fabricated with slots in their coating simulating craze cracks. These specimens were used to study oxidation damage detection and characterization using nondestructive evaluation (NDE) methods. These specimens were heat treated in air at 1143 and 1200 C to create cavities in the carbon substrate underneath the coating as oxygen reacted with the carbon and resulted in its consumption. The cavities varied in diameter from approximately 1 to 3 mm. Single-sided NDE methods were used since they might be practical for on-wing inspection, while x-ray micro-computed tomography (CT) was used to measure cavity sizes in order to validate oxidation models under development for carbon-carbon materials. An RCC sample having a naturally-cracked coating and subsequent oxidation damage was also studied with x-ray micro-CT. This effort is a follow-on study to one that characterized NDE methods for assessing oxidation damage in an RCC sample with drilled holes in the coating.

Roth, Don J.↗
NASA NTRS1993

Thermal certification tests of Orbiter Thermal Protection System tiles coated with KSC coating slurries

Thermal tests of Orbiter thermal protection system (TPS) tiles, which were coated with borosilicate glass slurries fabricated at Kennedy Space Center (KSC), were performed in the Radiant Heat Test Facility and the Atmospheric Reentry Materials & Structures Evaluation Facility at Johnson Space Center to verify tile coating integrity after exposure to multiple entry simulation cycles in both radiant and convective heating environments. Eight high temperature reusable surface insulation (HRSI) tiles and six low temperature reusable surface insulation (LRSI) tiles were subjected to 25 cycles of radiant heat at peaked surface temperatures of 2300 F and 1200 F, respectively. For the LRSI tiles, an additional cycle at peaked surface temperature of 2100 F was performed. There was no coating crack on any of the HRSI specimens. However, there were eight small coating cracks (less than 2 inches long) on two of the six LRSI tiles on the 26th cycle. There was practically no change on the surface reflectivity, physical dimensions, or weight of any of the test specimens. There was no observable thermal-chemical degradation of the coating either. For the convective heat test, eight HRSI tiles were tested for five cycles at a surface temperature of 2300 F. There was no thermal-induced coating crack on any of the test specimens, almost no change on the surface reflectivity, and no observable thermal-chemical degradation with an exception of minor slumping of the coating under painted TPS identification numbers. The tests demonstrated that KSC's TPS slurries and coating processes meet the Orbiter's thermal specification requirements.

Milhoan, James D.↗
NASA NTRS1998

Degradation of Teflon(tm) FEP Following Charged Particle Radiation and Rapid Thermal Cycling

During the Second Servicing Mission (SM2) of the Hubble Space Telescope (HST) severe degradation was observed on the outer layer of the thermal control blankets. Astronaut observations and photographs revealed large cracks in the metallized Teflon' FEP (fluorinated ethylene propylene), the outer layer of the multi-layer insulation (MLI), in many locations around the telescope. In an effort to understand what elements of the space environment might cause such damage, pristine Teflon(registered trademark) FEP was tested for durability to radiation and thermal cycling. Specimens were subjected to electron and proton fluences comparable to those experienced by HST and were subsequently thermal cycled in a custom-built rapid thermal cycle chamber. Tensile tests of the specimens showed that radiation followed by thermal cycling significantly reduced the ultimate strength and elongation of Teflon(registered trademark) FEP.

Townsend, Jacqueline A.↗
NASA NTRS1998

Degradation of Teflon(tm) FEP Following Charged Particle Radiation and Rapid Thermal Cycling

During the Second Servicing Mission (SM2) of the Hubble Space Telescope (HST) severe degradation was observed on the outer layer of the thermal control blankets. Astronaut observations and photographs revealed large cracks in the metallized Teflon(R) FEP (fluorinated ethylene propylene), the outer layer of the multi-layer insulation (MLI), in many locations around the telescope. In an effort to understand what elements of the space environment might cause such damage, pristine Teflon(R) FEP was tested for durability to radiation and thermal cycling. Specimens were subjected to electron and proton fluences comparable to those experienced by HST and were subsequently thermal cycled in a custom-built rapid thermal cycle chamber. Tensile tests of the specimens showed that radiation followed by thermal cycling significantly reduced the ultimate strength and elongation of Teflon(R) FEP.

Townsend, Jacqueline↗
NASA NTRS1999

Degradation of Teflon(trademark) FEP Following Charged Particle Radiation and Rapid Thermal Cycling

During the Second Servicing Mission (SM2) of the Hubble Space Telescope (HST) severe degradation was observed on the outer layer of the thermal control blankets. Astronaut observations and photographs revealed large cracks in the metallized Teflon(trademark) FEP (fluorinated ethylene propylene), the outer layer of the multi-layer insulation (MLI), in many locations around the telescope. In an effort to understand what elements of the space environment might cause such damage, pristine Teflon(trademark) FEP was tested for durability to radiation and thermal cycling. Specimens were subjected to electron and proton fluences comparable to those experienced by HST and were subsequently thermal cycled in a custom-built rapid thermal cycle chamber. Tensile tests of the specimens showed that radiation followed by thermal cycling significantly reduced the ultimate strength and elongation of Teflon(trademark) FEP.

Townsend, Jacqueline↗
NASA NTRS2000

Thermophysical and Thermomechanical Properties of Thermal Barrier Coating Systems

Thermal barrier coatings have been developed for advanced gas turbine and diesel engine applications to improve engine reliability and fuel efficiency. However, the issue of coating durability under high temperature cyclic conditions is still of major concern. The coating failure is closely related to thermal stresses and oxidation in the coating systems. Coating shrinkage cracking resulting from ceramic sintering and creep at high temperatures can further accelerate the coating failure process. The purpose of this paper is to address critical issues such as ceramic sintering and creep, thermal fatigue and their relevance to coating life prediction. Novel test approaches have been established to obtain critical thermophysical and thermomechanical properties of the coating systems under near-realistic temperature and stress gradients encountered in advanced engine systems. Emphasis is placed on the dynamic changes of the coating thermal conductivity and elastic modulus, fatigue and creep interactions, and resulting failure mechanisms during the simulated engine tests. Detailed experimental and modeling results describing processes occurring in the thermal barrier coating systems provide a framework for developing strategies to manage ceramic coating architecture, microstructure and properties.

Zhu, Dongming↗
NASA NTRS1989

Finite element analysis of the Space Shuttle Main Engine (SSME) high pressure fuel turbopump turbine blade (HPFTP)

Cracking of the turbine blades of the SSME HPFTP is studied using two- and three-dimensional finite element analysis. The development and composition of the two- and three-dimensional models are described. Analyses are conducted under the speed, pressure, and thermal load conditions that occur during the full power level of the engine. The effects of friction on the two-dimensional model are examined. The strain and life cycle data reveal that the LCF cracking in the first stage is not probable unless the effects of fit-up tolerance between the blade and rotor are present, and for the second stage it is predicted that hydrogen assisted LCF cracking will occur under the present thermal environment. Design modifications to alleviate this cracking are discussed.

Lee, H. M.↗
NASA NTRS2016

Thermal Cycling for Development of Hybrid Fuel for a Notional Mars Ascent Vehicle

A Mars Ascent Vehicle (MAV) would be a critical part of a potential Mars sample return campaign. As currently envisioned, the MAV would launch samples from Mars into Mars orbit. Challenging requirements would be placed on a MAV as it would need to be able to survive in the harsh Martian environment which can get as cold as -120°C in the winter. At these cold temperatures, survival heating could be necessary to keep the MAV within the allowable flight temperature range. Estimates for survival heating power show that a fuel capable of operating at low temperature would significantly reduce the total required energy of the MAV. This has led to consideration of a hybrid propulsion system due to its anticipated low temperature capability. A new wax-based fuel, SP7, has been developed for a notional MAV mission. Thermal modeling of the fuel in the Mars environment was completed in order to estimate the temperature response of the fuel during Mars diurnal temperature cycling. Preliminary thermal testing was conducted to determine the max temperature range, radial gradient, and ramp rate the fuel could tolerate without cracking. The modeling and preliminary testing was used to develop a comprehensive, 100-day thermal cycling campaign. A total of 200 cycles are planned that are representative of 200 Martian diurnal cycles, roughly one-third of a Martian year. After 150 cycles, the fuel samples show no signs of cracking that would jeopardize the rocket’s performance. Though not yet complete, this testing is showing that SP7 has the potential to successfully be used on a MAV mission.

Farias, Edgardo↗
NASA NTRS1975

Comparative thermal fatigue resistances of twenty-six nickel and cobalt base alloys

Thermal fatigue resistances were determined from fluidized bed tests. Cycles to cracking differed by almost three orders of magnitude for these materials with directional solidification and surface protection of definite benefit. The alloy-coating combination with the highest thermal fatigue resistance was directionally solidified NASA TAZ-8A with an RT-XP coating. It oxidation resistance was excellent, showing almost no weight change after 15 000 fluidized bed cycles.

Bizon, P. T.↗
NASA NTRS1992

Residual thermal and moisture influences on the analysis of local delaminations

An analysis using laminated plate theory is developed to calculate the strain energy release rate associated with local delaminations originating at off-axis, angle-ply, matrix cracks in laminates subjected to uniaxial loads. The analysis includes the contribution of residual thermal and moisture stresses to the strain energy released. Examples are calculated for the strain energy release rate associated with local delaminations originating at 90 deg and angle-ply(non 90 deg) matrix-ply cracks in glass-epoxy and graphite-epoxy laminates. The solution developed may be used to assess the relative contribution of mechanical, residual thermal, and moisture stresses on the strain energy release rate for local delamination for a variety of lay-ups and materials.

O'Brien, T. K.↗
NASA NTRS1985

Thermal fatigue tests of a radiative heat shield panel for a hypersonic transport

A pair of corrugation stiffened, beaded skin Rene 41 heat shield panels were exposed to 20,000 thermal cycles between room temperature and 1450 F to evaluate the thermal fatigue response of Rene 41 metallic heat shields for hypersonic cruise aircraft applications. At the conclusion of the tests, the panels retained substantial structural integrity; however, there were cracks and excessive wear in the vicinity of fastener holes and there was an 80-percent loss in ductility of the skin. Shrinkage of the panel which caused the cracks and wear must be considered in design of panels for Thermal Protection Systems (TPS) applications.

Webb, Granville L.↗
NASA NTRS1984

Ceramic composite liner material for gas turbine combustors

Advanced commercial and military gas turbine engines may operate at combustor outlet temperatures in excess of 1920 K (3000 F). At these temperatures combustors liners experience extreme convective and radiative heat fluxes. The ability of a plasma sprayed ceramic coating to reduce liner metal temperature has been recognized. However, the brittleness of the ceramic layer and the difference in thermal expansion with the metal substrate has caused cracking, spalling and some separation of the ceramic coating. Research directed at turbine tip seals (or shrouds) has shown the advantage of applying the ceramic to a compliant metal pad. This paper discusses recent studies of applying ceramics to combustor liners in which yttria stabilized zirconia plasma sprayed on compliant metal substrates which were exposed to near stoichiometric combustion, presents performance and durability results, and describes a conceptual design for an advanced, small gas turbine combustor. Test specimens were convectively cooled or convective-transpiration cooled and were evaluated in a 10 cm square flame tube combustor at inlet air temperatures of 533 K (500 F) and at a pressure of 0.5 MPa (75 psia). The ceramics were exposed to flame temperatures in excess of 2000 K (3320 F). Results appear very promising with all 30 specimens surviving a screening test and one of two specimens surviving a cyclic durability test.

Ercegovic, D. B.↗
NASA NTRS1986

Fast Glazing of Alumina/Silica Tiles

Technique for applying ceramic coating to fibrous silica/alumina insulation tiles prevents cracks and substantially reduces firing time. To reduce thermal stresses in tile being coated, high-temperature, shorttime firing schedule implemented. Such schedule allows coating to mature while substrate remains at relatively low temperature, reducing stress differential between coating and substrate. Technique used to repair tiles with damaged coatings and possibly used in heat-treating objects made of materials having different thermal-expansion coefficients.

Creedon, J. F.↗