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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS2001

Damage Assessment of Creep Tested and Thermally Aged Udimet 520 Using Acousto-Ultrasonics

Due to elevated temperatures and excessive stresses, turbine components may experience creep behavior. As a result, it is desirable to monitor and assess the current condition of such components. This study employed the Acousto-Ultrasonics (AU) method in an effort to monitor the state of the material at various percentages of used up creep life in the nickel base alloy, Udimet 520. A stepped specimen (i.e., varying cross sectional area) was employed which allowed for a postmortem nondestructive evaluation (NDE) analysis of the various levels of used up life. The overall objectives here were two fold: First, a user friendly, graphical interface AU system was developed, and second the new AU system was applied as an NDE tool to assess distributed damage resulting from creep. The experimental results demonstrated that the AU method shows promise as an NDE tool capable of detecting material changes as a function of used up creep life. Furthermore, the changes in the AU parameters were mainly attributed to the case of combined load and elevated temperature (i.e., creep) and not simply because of a timed exposure at elevated temperature (i.e., heat treatment or thermal aging).

Gyekenyesi, Andrew L.↗
NASA NTRS1998

On-Orbit Teflon(trademark) FEP Degradation

During the Hubble Space Telescope (HST) Second Servicing Mission (SM2), degradation of unsupported Teflon(trademark) FEP (fluorinated ethylene propylene), used as the outer layer of the multi-layer insulation (MLI) blankets, was evident as large cracks on the telescope light shield. A sample of the degraded outer layer was retrieved during the mission and returned to Earth for ground testing and evaluation. The results of the Teflon(trademark) FEP sample evaluation and additional testing of pristine Teflon(trademark) FEP led the investigative team to theorize that the HST damage was caused by thermal cycling with deep-layer damage from electron and proton radiation which allowed the propagation of cracks along stress concentrations, and that the damage increased with the combined total dose of electrons, protons, UV and x-rays along with thermal cycling. This paper discusses the testing and evaluation of the Teflon(trademark) FEP.

Townsend, Jacqueline A.↗
NASA NTRS1998

On-Orbit Teflon FEP Degradation

During the Hubble Space Telescope (HST) Second Servicing Mission (SM2), degradation of unsupported Teflon' FEP (fluorinated ethylene propylene), used as the outer layer of the multi-layer insulation (MLI) blankets, was evident as large cracks on the telescope light shield. A sample of the degraded outer layer was retrieved during the mission and returned to Earth for ground testing and evaluation. The results of the Teflon FEP sample evaluation and additional testing of pristine Teflon FEP led the investigative team to theorize that the HST damage was caused by thermal cycling with deep-layer damage from electron and proton radiation which allowed the propagation of cracks along stress concentrations, and that the damage increased with the combined total dose of electrons, protons, LTV and x-rays along with thermal cycling. This paper discusses the testing and evaluation of the Teflon FEP.

Townsend, Jacqueline A.↗
NASA NTRS1998

On-Orbit Teflon(R) FEP Degradation

During the Hubble Space Telescope (HST) Second Servicing Mission (SM2), degradation of unsupported Teflon(R) FEP (fluorinated ethylene propylene), used as the outer layer of the multi-layer insulation (MLI) blankets, was evident as large cracks on the telescope light shield. A sample of the degraded outer layer was retrieved during the mission and returned to Earth for ground testing and evaluation. The results of the Teflon(R) FEP sample evaluation and additional testing of pristine Teflon(R) FEP led the investigative team to theorize that the HST damage was caused by thermal cycling with deep-layer damage from electron and proton radiation which allowed the propagation of cracks along stress concentrations, and that the damage increased with the combined total dose of electrons, protons, UV and x-rays along with thermal cycling. This paper discusses the testing and evaluation of the Teflon(R) FEP.

Townsend, Jacqueline A.↗
NASA NTRS1999

On-Orbit Teflon(trademark) FEP Degradation

During the Hubble Space Telescope (HST) Second Servicing Mission (SM2), degradation of unsupported Teflon(trademark) FEP (fluorinated ethylene propylene), used as the outer layer of the multi-layer insulation (MLI) blankets, was evident as large cracks on the telescope light shield. A sample of the degraded outer layer was retrieved during the mission and returned to Earth for ground testing and evaluation. The results of the Teflon(trademark) FEP sample evaluation and additional testing of pristine Teflon FEP led the investigative team to theorize that the HST damage was caused by thermal cycling with deep-layer damage from electron and proton radiation which allowed the propagation of cracks along stress concentrations, and that the damage increased with the combined total dose of electrons, protons, UV and x-rays along with thermal cycling. This paper discusses the testing and evaluation of the Teflon(trademark) FEP.

Townsend, Jacqueline↗
NASA NTRS1983

A conceptual model of the thermal-spike mechanism in graphte/epoxy laminates

The influence of a thermal spike on a moisture-saturated graphite/epoxy composite was studied in detail. A single thermal spike from 25 to 132 C was found to produce damage, as evidenced by a significant increase in the level of moisture saturation in the composite. Approximately half of this increase remained after a vacuum-anneal at 150 C for seven days, suggesting the presence of an irreversible damage component. Subsequent thermal spikes created less and less additional moisture absorption, with the cumulative effect being a maximum or limiting moisture capacity of the composite. These observations are explained in terms of a model previously developed to explain the reverse thermal effect of moisture absorption in epoxy and epoxy-matrix composites. This model, based on the inverse temperature dependence of free volume, improves current understanding of thermal spike effects in graphite/epoxy composites.

Adamson, M. J.↗
NASA NTRS2006

Orbiter Return-To-Flight Entry Aeroheating

The Columbia accident on February 1, 2003 began an unprecedented level of effort within the hypersonic aerothermodynamic community to support the Space Shuttle Program. During the approximately six month time frame of the primary Columbia Accident Investigation Board activity, many technical disciplines were involved in a concerted effort to reconstruct the last moments of the Columbia and her crew, and understand the critical events that led to that loss. Significant contributions to the CAIB activity were made by the hypersonic aerothermodynamic community(REF CAIB) in understanding the re-entry environments that led to the propagation of an ascent foam induced wing leading edge damage to a subsequent breech of the wing spar of Columbia, and the subsequent breakup of the vehicle. A core of the NASA hypersonic aerothermodynamics team that was involved in the CAIB investigation has been combined with the United Space Alliance and Boeing Orbiter engineering team in order to position the Space Shuttle Program with a process to perform in-flight Thermal Protection System damage assessments. This damage assessment process is now part of the baselined plan for Shuttle support, and is a direct out-growth of the Columbia accident and NASAs response. Multiple re-entry aeroheating tools are involved in this damage assessment process, many of which have been developed during the Return To Flight activity. In addition, because these aeroheating tools are part of an overall damage assessment process that also involves the thermal and stress analyses community, in addition to a much broader mission support team, an integrated process for performing the damage assessment activities has been developed by the Space Shuttle Program and the Orbiter engineering community. Several subsets of activity in the Orbiter aeroheating communities support to the Return To Flight effort have been described in previous publications (CFD?, Cavity Heating? Any BLT? Grid Generation?). This work will provide a description of the integrated process utilized to perform Orbiter tile damage assessment, and in particular will seek to provide a description of the integrated aeroheating tools utilized to perform these assessments. Individual aeroheating tools will be described which provide the nominal re-entry heating environment characterization for the Orbiter, the heating environments for tile damage, heating effects due to exposed Thermal Protection System substrates, the application of Computational Fluid Dynamics for the description of tile cavity heating, and boundary layer transition prediction. This paper is meant to provide an overall view of the integrated aeroheating assessment process for tile damage assessment as one of a sequence of papers on the development of the boundary layer transition prediction capability in support of Space Shuttle Return To Flight efforts.

Campbell, Charles H.↗
NASA NTRS1972

Influence of creep damage on the low cycle thermal-mechanical fatigue behavior of two tantalum base alloys

Low cycle fatigue tests have been performed on the tantalum base alloys T-111 and ASTAR 811C with synchronized, independently programmed temperature and strain cycling. The thermal-mechanical cycles applied fell into three basic categories: these were isothermal cycling, in-phase thermal cycling, and out-of-phase thermal cycling. In-phase cycling was defined as tensile deformation associated with high temperature and compressive deformation with low temperature, while out-of-phase thermal cycling was defined as the reverse case. The in-phase thermal cycling had a pronounced detrimental influence on the fatigue life of both alloys, with the life reduction being greater in the solid solution strengthened T-111 alloy than in the carbide strengthened ASTAR 811C alloy. The out-of-phase tests also showed pronounced effects on the fatigue life of both alloys, although not as dramatic.

Sheffler, K. D.↗
NASA NTRS1982

A model of the thermal-spike mechanism in graphite/epoxy laminates

The influence of a thermal spike on a moisture-saturated graphite/epoxy composite was studied in detail. A single thermal spike from 25 C to 132 C was found to produce damage as evidenced by a significant increase in the level of moisture saturation in the composite. Approximately half of this increase remained after a vacuum anneal at 150 C for 7 days, suggesting the presence of an irreversible damage component. Subsequent thermal spikes created less and less additional moisture absorption, with the cumulative effect being a maximum or limiting moisture capacity of the composite. These observations are explained in terms of a model previously developed to explain the reverse thermal effect of moisture absorption in epoxy and epoxy matrix composites. This model, based on the inverse temperature dependence of free volume, contributes an improved understanding of thermal-spike effects in graphite/epoxy composites.

Adamson, M. J.↗
NASA NTRS1991

MLITemp: A computer program to predict the thermal effects associated with hypervelocity impact damage to space station MLI

A family of user-friendly, DOS PC based, Microsoft BASIC programs written to provide spacecraft designers with empirical predictions of space debris damage to orbiting spacecraft are described. Spacecraft wall temperatures and condensate formation is also predicted. The spacecraft wall configuration is assumed to consist of multilayered insulation (MLI) placed between a Whipple style bumper and the pressure wall. Impact damage predictions are based on data sets of experimental results obtained from simulating debris impacts on spacecraft using light gas guns on earth. A module of the program facilitates the creation of the database of experimental results that is used by the damage prediction modules to predict damage to the bumper, the MLI, and the pressure wall. A finite difference technique is used to predict temperature distributions in the pressure wall, the MLI, and the bumper. Condensate layer thickness is predicted for the case where the pressure wall temperature drops below the dew point temperature of the spacecraft atmosphere.

Rule, W. K.↗
NASA NTRS1978

Annealing of GaAs solar cells damaged by electron irradiation

Measurements of thermal annealing of GaAlAs/GaAs solar cells damaged by 1 MeV electron irradiation are reported, and the magnitude of the short-circuit current recovery is discussed. The damaged cells are annealed in a vacuum at 200 C. A cell irradiated at 10 to the 13th power electrons per sq cm recovers all its lost short-circuit current after 15 hours of annealing. Possible application of the annealing process to solar cells in space is also considered.

Walker, G. H.↗
NASA NTRS2009

Mid-Infrared Reflectance Imaging of Thermal-Barrier Coatings

An apparatus for mid-infrared reflectance imaging has been developed as means of inspecting for subsurface damage in thermal-barrier coatings (TBCs). The apparatus is designed, more specifically, for imaging the progression of buried delamination cracks in plasma-sprayed yttria-stabilized zirconia coatings on turbine-engine components. Progression of TBC delamination occurs by the formation of buried cracks that grow and then link together to produce eventual TBC spallation. The mid-infrared reflectance imaging system described here makes it possible to see delamination progression that is invisible to the unaided eye, and therefore give sufficiently advanced warning before delamination progression adversely affects engine performance and safety. The apparatus (see figure) includes a commercial mid-infrared camera that contains a liquid-nitrogen-cooled focal plane indium antimonide photodetector array, and imaging is restricted by a narrow bandpass centered at wavelength of 4 microns. This narrow wavelength range centered at 4 microns was chosen because (1) it enables avoidance of interfering absorptions by atmospheric OH and CO2 at 3 and 4.25 microns, respectively; and (2) the coating material exhibits maximum transparency in this wavelength range. Delamination contrast is produced in the midinfrared reflectance images because the introduction of cracks into the TBC creates an internal TBC/air-gap interface with a high diffuse reflectivity of 0.81, resulting in substantially higher reflectance of mid-infrared radiation in regions that contain buried delamination cracks. The camera is positioned a short distance (.12 cm) from the specimen. The mid-infrared illumination is generated by a 50-watt silicon carbide source positioned to the side of the mid-infrared camera, and the illumination is collimated and reflected onto the specimen by a 6.35-cm-diameter off-axis paraboloidal mirror. Because the collected images are of a steady-state reflected intensity (in contrast to the transient thermal response observed in infrared thermography), collection times can be lengthened to whatever extent needed to achieve desired signal-to-noise ratios. Each image is digitized, and the resulting data are processed in several steps to obtain a true mid-infrared reflectance image. The raw image includes thermal radiation emitted by the specimen in addition to the desired reflected radiation. The thermal-radiation contribution is eliminated by subtracting the image obtained with the illumination off from the image obtained with the illumination on. A flat-field correction is then made to remove the effects of non-uniformities in the illumination level and pixel-to-pixel variations in sensitivity. This correction is performed by normalizing to an image of a standard object that has a known reflectance at a wavelength of 4 microns. After correction, each pixel value is proportional to the reflectance (at a wavelength of 4-microns) at the corresponding location on the specimen. Mid-infrared reflectance imaging of specimens that were thermally cyc led for different numbers of cycles was performed and demonstrated t hat mid-infrared reflectance imaging was able to monitor the gradual delamination progression that occurs with continued thermal cycling . Reproducible values were obtained for the reflectance associated w ith an attached and fully delaminated TBC, so that intermediate refle ctance values could be interpreted to successfully predict the numbe r of thermal cycles to failure.

Edlridge, Jeffrey I.↗
NASA NTRS1991

Composite material systems for hydrogen management

The task of managing hydrogen entry into elevated temperature structural materials employed in turbomachinery is a critical engineering area for propulsion systems employing hydrogen or decomposable hydrocarbons as fuel. Extant structural materials, such as the Inconel series, are embrittled by the ingress of hydrogen in service, leading to a loss of endurance and general deterioration of load-bearing dependability. Although the development of hydrogen-insensitive material systems is an obvious engineering option, to date insensitive systems cannot meet the time-temperature-loading service extremes encountered. A short-term approach that is both feasible and technologically sound is the development and employment of hydrogen barrier coatings. The present project is concerned with developing, analyzing, and physically testing laminate composite hydrogen barrier systems, employing Inconel 718 as the structural material to be protected. Barrier systems will include all metallic, metallic-to-ceramic, and, eventually, metallic/ceramic composites as the lamellae. Since space propulsion implies repetitive engine firings without earth-based inspection and repair, coating durability will be closely examined, and testing regimes will include repetitive thermal cycling to simulate damage accumulation. The target accomplishments include: generation of actual hydrogen permeation data for metallic, ceramic-metallic, and hybrid metallic/ceramic composition barrier systems, practically none of which is currently extant; definition of physical damage modes imported to barrier systems due to thermal cycling, both transient temperature profiles and steady-state thermal mismatch stress states being examined as sources of damage; and computational models that incorporate general laminate schemes as described above, including manufacturing realities such as porosity, and whatever defects are introduced through service and characterized during the experimental programs.

Pangborn, R. N.↗
NASA NTRS2014

Monitoring of Thermal Protection Systems and MMOD using Robust Self-Organizing Optical Fiber Sensing Networks

The general aim of this work is to develop and demonstrate a prototype structural health monitoring system for thermal protection systems that incorporates piezoelectric acoustic emission (AE) sensors to detect the occurrence and location of damaging impacts, such as those from Micrometeoroid Orbital Debris (MMOD). The approach uses an optical fiber Bragg grating (FBG) sensor network to evaluate the effect of detected damage on the thermal conductivity of the TPS material. Following detection of an impact, the TPS would be exposed to a heat source, possibly the sun, and the temperature distribution on the inner surface in the vicinity of the impact measured by the FBG network. A similar procedure could also be carried out as a screening test immediately prior to re-entry. The implications of any detected anomalies in the measured temperature distribution will be evaluated for their significance in relation to the performance of the TPS during reentry. Such a robust TPS health monitoring system would ensure overall crew safety throughout the mission, especially during reentry.

Fiber Optic Sensors↗