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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 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

Creep Testing of High-Temperature Cu-8 Cr-4 Nb Alloy Completed

A Cu-8 at.% Cr-4 at.% Nb (Cu-8 Cr-4 Nb) alloy is under development for high-temperature, high heatflux applications, such as actively cooled, hypersonic vehicle heat exchangers and rocket engine combustion chambers. Cu-8 Cr-4 Nb offers a superior combination of strength and conductivity. It has also shown exceptional low-cycle fatigue properties. Following preliminary testing to determine the best processing route, a more detailed testing program was initiated to determine the creep lives and creep rates of Cu-8 Cr-4 Nb alloy specimens produced by extrusion. Testing was conducted at the NASA Lewis Research Center with constant-load vacuum creep units. Considering expected operating temperatures and mission lives, we developed a test matrix to accurately determine the creep properties of Cu-8 Cr-4 Nb between 500 and 800 C. Six bars of Cu-8 Cr-4 Nb were extruded. From these bars, 54 creep samples were machined and tested. The figure on the left shows the steady-state, or second-stage, creep rates for the samples. Comparison data for NARloy-Z (Cu-3 wt % Ag-0.5 wt % Zr), the alloy currently used in combustion chamber liners, were not unavailable. Therefore the steady-state creep rates for Cu at similar temperatures are presented. As expected, in comparison to pure Cu, the creep rates for Cu-8 Cr-4 Nb are much lower. The lives of the samples are presented in the figure on the right. As shown, Cu-8 Cr-4 Nb at 800 C is comparable to NARloy-Z at 648 C. At equivalent temperatures, Cu-8 Cr-4 Nb enjoys a 20 to 50 percent advantage in stress for a given life and 1 to 3 orders of magnitude greater life at a given stress. The improved properties allow for design tradeoffs and improvements in new and existing heat exchangers such as the next generation of combustion chamber liners. Average creep rates for Cu-8 Cr-4 Nb and pure Cu are shown. Average creep lives for Cu-8 Cr- 4 Nb and NARloy-Z are also shown. Currently, two companies are interested in the commercial usage of the Cu-8 Cr-4 Nb alloy. The Rocketdyne Division of Rockwell International is conducting independent testing to analyze the properties for their projected needs in advanced rocket engine applications. Metallamics, a company based in Traverse City, Michigan, is entering into a Space Act Agreement to evaluate and test Cu-Cr-Nb alloys as materials for welding electrodes that are used in robotic welding operations. Creep rate is one of the alloy properties that determines the degree to which a welding electrode will mushroom or expand at the tip. A material with a low creep rate will resist mushrooming and give the electrode a longer life, minimizing downtime. This application holds the potential for large-scale usage of the alloy in the automotive and other industries. Success here would dramatically decrease the cost of the alloy and increase availability for aerospace applications.

Source record↗
NASA NTRS2005

Forging of Advanced Disk Alloy LSHR

The powder metallurgy disk alloy LSHR was designed with a relatively low gamma precipitate solvus temperature and high refractory element content to allow versatile heat treatment processing combined with high tensile, creep and fatigue properties. Grain size can be chiefly controlled through proper selection of solution heat treatment temperatures relative to the gamma precipitate solvus temperature. However, forging process conditions can also significantly influence solution heat treatment-grain size response. Therefore, it is necessary to understand the relationships between forging process conditions and the eventual grain size of solution heat treated material. A series of forging experiments were performed with subsequent subsolvus and supersolvus heat treatments, in search of suitable forging conditions for producing uniform fine grain and coarse grain microstructures. Subsolvus, supersolvus, and combined subsolvus plus supersolvus heat treatments were then applied. Forging and subsequent heat treatment conditions were identified allowing uniform fine and coarse grain microstructures.

Gabb, Timothy P.↗
NASA NTRS2004

Coating Development for GRCop-84 Liners for Reusable Launch Vehicles Aided by Modeling Studies

The design of the next generation of reusable launch vehicles calls for using GRCop-84 copper alloy liners based on a composition invented at the NASA Glenn Research Center. Despite its considerable advantage over other copper alloys, it is expected that GRCop-84 will suffer from environmental degradation depending on the type of rocket fuels used and on thermomechanical fatigue. Applying protective coatings on GRCop-84 substrates can minimize or eliminate many of these problems and extend the operational life of the combustion liner. This could increase component reliability, shorten depot maintenance turnaround times, and lower operating costs. Therefore, Glenn is actively pursuing the development of advanced coatings technology for GRCop-84 liners. Technology is being developed in four major areas: (1) new metallic coating compositions, (2) application techniques, (3) test methods, and (4) life prediction design methodology using finite element analysis. The role of finite element analysis in guiding the coating effort is discussed in this report. Thermal analyses were performed at Glenn for different combinations of top- and bondcoat compositions to determine the temperature variation across the coated cross section with the thickness of the top coat. These calculations were conducted for simulated LH2/LO2 booster engine conditions assuming that the bond coat had a constant thickness of 50 m. The preceding graphs show the predicted temperatures at the outer surface of the top coat (hot wall), at the top-coat/bond-coat interface, at the bond-coat/GRCop-84 interface, and at the GRCop-84 cold wall as a function of top-coat thickness for Cu- 26(wt%)Cr top coat (top graph), Ni-17(wt%)Cr-6%Al-0.5%Y top coat and Cu-26%Cr bond coat, and NiAl top coat and Ni bond coat. In all cases, the temperature of the top coat at the hot wall increased with increasing top-coat thickness and with corresponding decreases in the temperatures at the two interfaces and the cold wall. These temperatures are not acutely sensitive to the thermal conductivity of the top coat when it exceeds 25 and 50 W/m/K for low and high heat flux engines. This observation is significant for two reasons. First, several different top-coat compositions can be evaluated as potential protective coatings without loss in the heat-transfer efficiency of the coated system. Second, materials with thermal conductivities less than the critical values of 25 or 50 W/m/K are more likely to act as thermal barrier coatings. The deposition of overlay coatings on GRCop-84 substrates results in the development of residual stresses. The presence of these residual stresses influences the probability of coating spallation, the thermal cycling life, and the fatigue properties of the coated substrate during use. Since it is important to understand how these stresses develop during the vacuum-plasma-spraying coating deposition process, the nature and magnitudes of the cool-down residual stresses were calculated and compared with experimentally determined values across the coated cross section of a disk specimen. The calculations were conducted assuming that the specimen cools down to room temperature from vacuum plasma-spraying temperatures of either 250 or 650 C. The effects of coating the substrate with and without grit blasting were also theoretically examined. The final graph compares the predicted and the experimental results for a GRCop-84 disk coated with about a 50- m-thick Ni bond coat and a 75- to 100- m NiAl top coat, where the curves for NASA-2 assume the presence of a prior residual stress generated by grit blasting under conditions similar to the experimental situation. The predicted cool-down in-plane stresses were compressive in both the NiAl top coat and the Ni bond coat. They were also compressive in the substrate to a depth of about 0.25 mm from the Ni/GRCop-84 interface when the vacuum-plasma-spraying temperature was low. However, using a higher plasma spraying temperaturs likely to leave the substrate under a small tensile stress to counter the compressive stresses in the bond and top coats because of the relaxation of residual stresses generated in the substrate during the grit blasting of its surface prior to spraying. These results suggest that the NiAl and Ni coatings are unlikely to spall after spraying as confirmed by the microstructural observations shown in the following photomicrograph of an as-sprayed specimen. Finally, it is noted that the calculated and experimental results are not in complete agreement, which indicates that both the experimental and modeling techniques need further refinement.

Raj, Sai V.↗
NASA NTRS2006

Effects of Hydrogen on GRCop-84

This report is a section of the final report on the GRCop-84 task of the Constellation Program and incorporates the results obtained between October 2000 and September 2005 when the program ended. GRCop-84 contains approximately 5.5 wt% Nb. Nb can react with H and embrittle easily. Previous work had indicated the thermodynamic possibility that Cr 2 Nb could react with H and form niobium hydrides in the presence of high pressure H such as seen in the Space Shuttle Main Engine. In this study, samples were charged with H and then tested in both high pressure H and He environments to determine if measurable differences existed which indicate that hydrogen embrittlement occurs in GRCop-84. Tensile, notched tensile, stress rupture and low cycle fatigue properties were surveyed. High pressure H environment stress rupture testing resulted in a lower reduction in area than a high pressure He environment, and the LCF lives at high strain ranges fall below the lower 95 percent confidence interval for the baseline data, but in general no significant differences were noted either between H and He environment tests or between hydrogen charged materials and the baseline, uncharged extruded GRCop-84 data sets. There was also no discernable evidence of the formation of hydrides or changes in fracture morphology indicating hydrogen embrittlement occurred.

Ellis, David L.↗
NASA NTRS2010

Experimental Evaluation of Fatigue Crack Initiation from Corroded Hemispherical Notches in Aerospace Structural Materials

A test program was developed and executed to evaluate the influence of corroded hemispherical notches on the fatigue crack initiation and propagation in aluminum 7075-T7351, 4340 steel, and D6AC steel. Surface enhancements such as shot peening and laser shock peening were also incorporated as part of the test effort with the intent of improving fatigue performance. In addition to the testing, fracture mechanics and endurance limit based analysis methods were evaluated to characterize the results with the objective of challenging typical assumptions used in modeling fatigue cracks from corrosion pits. The results specifically demonstrate that the aluminum and steel alloys behave differently with respect to fatigue crack initiation from hemispherical corrosion pits. The aluminum test results were bounded by the fracture mechanics and endurance limit models while exhibiting a general insensitivity to the residual stress field generated by shot peening. The steel specimens were better characterized by the endurance limit fatigue properties and did exhibit sensitivities to residual stresses from the shot peening and laser shock peening

Garcia, Daniel B.↗
NASA NTRSDate not supplied

Porosity Prediction and Detection During Composite Cure Using Simulation and Ultrasonic In-Situ Inspection Inside and Autoclave

Composite materials offer unique benefits in aerospace applications such as increased strength-to-weight ratio and improved fatigue properties. They are increasingly being used in major commercial aircraft programs. However, current processing methods can lead to defects in composite parts, which are currently identified using post-manufacturing inspection methods. A cure defects process model has been developed to predict the formation of manufacturing defects (e.g., porosity and fiber waviness) in composites based on the cure parameters and part geometry. However, a capability to directly validate porosity during cure did not exist. Validation methods included comparing resin pressure measured during cure with predictions by the process model and inspection/microscopy after cure. This study developed a high-temperature ultrasonic inspection system to detect porosity defects in composites during autoclave cure and experimentally verify the predictions of the model. The system operated inside an autoclave within an enclosure cooled by intermittent liquid nitrogen(LN2) injections. A high-temperature 2.25 MHz ultrasonic transducer was utilized to transmit ultrasonic waves through the tool plate and into the composite material and to measure the amplitude and time of flight of the reflected waves with a 1 mm × 1 mm resolution. Porosity was observed via the ultrasonic reflections, which experienced increased attenuation in regions of high porosity. Distinct regions of increased porosity were present due to uneven pressure across the panel, which was driven by an intentional misfit between the flat caul plate and the tapered composite panel with ply drops. These observations matched the predictions of the model using the inputs from the experiment (e.g., part material and geometry, cure cycle) and were validated by post-cure ultrasonic inspection and micrographs. The in-situ inspection system was able to successfully provide defect detection and localization and can be applied to future manufacturing of composite structure for aerospace applications.

Porosity↗
NASA NTRS2021

Machine Learning to Predict Joint Performance in Epoxy Composites Based on Process Parameters

Polymer matrix composites are gaining popularity in the aerospace industry due to their high specific strength, fatigue properties, and processability. However, based on current FAA certification guidelines, manufacturers utilizing current state-of-the-art composites made with adhesive bonds commonly install redundant fasteners to guarantee the strength of these adhesively bonded composite parts. The number of fasteners in a single-aisle commercial transport aircraft is typically on the order of 105, which reduces manufacturing rate, increases cost tremendously, and reduces the advantage of the specific strength composites provide. Due to this, the Adhesive Free Bonding of Composites (AERoBOND) project at NASA Langley Research Center has developed a novel assembly process to manufacture complex composite parts without the use of adhesives and fasteners. However, optimization of the process is currently challenging due to the complex and interdependent process parameters. To assist with the optimization, four machine learning algorithms utilizing gradient boosting decision trees were created to provide predictions for the mechanical and characterization properties of the composite parts. Approximately 200 random states from each algorithm were tested, and the models from each state were isolated and analyzed based on their accuracy, a validation process, and their feature importance. This analysis concluded that the models created from the machine learning algorithms could accelerate a parametric study for the AERoBOND process by rapidly optimizing process parameters to achieve desired performance characteristics.

Brennen Michael Middleton↗
NASA NTRSDate not supplied

Machine Learning to Predict Joint Performance in Epoxy Composites Based on Process Parameters

Polymer matrix composites are gaining popularity in the aerospace industry due to their high specific strength, fatigue properties, and processability. However, based on current FAA certification guidelines, manufacturers utilizing current state-of-the art composites made with adhesive bonds commonly install redundant fasteners to guarantee the strength of these adhesively bonded composite parts.1,2 The number of fasteners in a single-aisle commercial transport aircraft is typically on the order of 105, which reduces manufacturing rate, increases cost tremendously, and reduces the advantage of the specific strength composites provide. Due to this, the Adhesive Free Bonding of Composites (AERoBOND) project at NASA Langley Research Center has developed a novel assembly process to manufacture complex composite parts without the use of adhesives and fasteners.1 However, optimization of the process is currently challenging due to the complex and interdependent process parameters. To assist with the optimization, four machine learning algorithms utilizing gradient boosting decision trees were created to provide predictions for the mechanical and characterization properties of the composite parts. Approximately 200 random states from each algorithm were tested, and the models from each state were isolated and analyzed based on their accuracy, a validation process, and their feature importance. This analysis concluded that the models created from the machine learning algorithms could accelerate a parametric study for the AERoBOND process by rapidly optimizing process parameters to achieve desired performance characteristics.

Brennen M Middleton↗
NASA NTRS2021

Porosity Detection and Localization During Composite Cure Inside an Autoclave Using Ultrasonic Inspection

Composite materials offer unique benefits in aerospace applications such as increased strength-to-weight ratio and improved fatigue properties. They are increasingly being used in major commercial aircraft programs. However, current processing methods can lead to defects in composite parts, which must be detected using post-manufacturing inspection methods. Porosity (i.e., pores, voids) is a critical defect from the cure process that is detrimental to performance of the composite part. It is therefore necessary to understand and eliminate the formation of porosity defects. At NASA Langley Research Center, an in-situ cure monitoring system was developed to detect porosity defects as they form in real-time inside an autoclave. A capability to directly detect and localize porosity within a composite during cure did not exist before. This study is focused on an elevated-temperature ultrasonic inspection system to detect porosity defects in composites during autoclave cure. The ultrasonic inspection system operated inside an autoclave within an enclosure cooled by intermittent liquid nitrogen (LN_2) injections. A high-temperature, 2.25 MHz, transducer transmitted ultrasonic waves through the tool plate and into the composite part and measured the amplitude and time of flight of the reflected waves with a 1 mm × 1 mm step size/areal resolution. Porosity was observed via the ultrasonic reflections, which experienced increased attenuation in regions of high porosity. Distinct regions of increased porosity were present due to uneven pressure across the panel, which was driven by an intentional misfit between the flat caul plate and the tapered composite panel with ply drops. The results were validated by post-cure ultrasonic inspection and micrographs. The in-situ inspection system was able to successfully provide porosity detection and localization in the tapered ply-drop panel. The successful results indicate the promise of this system for future implementation in the manufacturing of composite structures for aerospace applications.

Composites↗
NASA NTRSDate not supplied

ASTM International Conference on Additive Manufacturing – 2022

The In-Situ Project: Correlating In-situ monitoring data and Non-Destructive Evaluation (NDE) methods to characterize defect populations in Laser Powder Bed Fusion (L-PBF) material NDE is critical to enable the safe use of additive manufactured (AM) parts in aerospace applications. However, traditional NDE methods are not fully adaptable to complex AM geometries to appropriately screen for critical defects, and L-PBF processes inherently generate defects. In-situ monitoring techniques can give information about the quality of the build process and indicate process variations that could potentially create defects. Proven, causal correlations between in-situ monitoring indications and resulting defects are needed to inform about the final quality of the part. Thus, the development of a new approach is necessary. NASA started a two-phase study that would investigate techniques for characterizing defect populations in L-PBF material, first in material with induced defects, then in a nominal L-PBF process, by correlating in-situ, NDE, and serial sectioning data. Objectives also include the evaluation of seeded defect methodologies for realistic defect creation, creating a baseline of defect populations in a Qualified Material Process (QMP), characterizing the effect of heat treatment, and correlating defect populations with tensile and fatigue properties. By comparing computed tomography images, in-situ monitoring images, and metallography images, the team has been able to draw preliminary conclusions. They include a better understanding of the limitation of the NDE tools, learning about the “healing” of the samples observed during layer building and a comparison study on different laser power variation samples. Those preliminary conclusions will help dictate the follow-up work that will be presented at the conference.

additive manufacturing↗
NASA NTRS1985

Effect of low temperature on fatigue and fracture properties of Ti-5Al-2.5Sn(ELI) for use in engine components

Experiments were conducted to evaluate the characteristics of the Ti-5Al-2.5Sn (ELI) alloy used in a fuel pump impeller at cryogenic temperatures. Tension, fracture toughness, and fatigue crack propagation data were collected determining the effect of frequency and load ratio on crack propagation. The results revealed that tensile strength increased significantly at 20 K compared to room temperature and fracture toughness was reduced at cryogenic temperatures. The fatigue crack growth rate was not sensitive to experimental conditions and there were only minimal crack orientation effects. Different frequencies produced no effect. At various temperatures and frequencies a load ratio increase resulted in higher crack growth rates. At low stress intensity levels the fatigue rate for both temperatures was the same; however, at high stress intensity levels the crack growth rate at 20 K increased because of the decrease in fracture toughness. The results correlated well with previous data.

Ryder, J. T.↗
NASA NTRS1976

Fatigue of notched fiber composite laminates. Part 2: Analytical and experimental evaluation

The analytical/experimental correlation study was performed to develop an understanding of the behavior of notched Boron/epoxy laminates subjected to tension/tension fatigue loading. It is postulated that the fatigue induced property changes (stiffness as well as strength) of the laminate can be obtained from the lamina fatigue properties. To that end, the Boron/epoxy lamina static and fatigue data (lifetime, residual stiffness and strength) were obtained initially. The longitudinal and transverse tension data were determined from the (0) and (90) laminate tests while the in-plane shear data were obtained from the (+ or - 45) sub s laminates. The static tests obtained the notched strength and mode of failure while the fatigue tests determined lifetime, damage propagation and residual strength. The failure in static tension occurred in a transverse crack propagation mode.

Kulkarni, S. V.↗
NASA NTRS1976

YF-12 Lockalloy ventral fin program, volume 1

Results are presented of the YF-12 Lockalloy Ventral Fin Program which was carried out by Lockheed Aircraft Corporation - Advanced Development Projects for the joint NASA/USAF YF-12 Project. The primary purpose of the program was to redesign and fabricate the ventral fin of the YF-12 research airplane (to reduce flutter) using Lockalloy, and alloy of beryllium and aluminum, as a major structural material. A secondary purpose, was to make a material characterization study (thermodynamic properties, corrosion; fatigue tests, mechanical properties) of Lockalloy to validate the design of the ventral fin and expand the existing data base on this material. All significant information pertinent to the design and fabrication of the ventral fin is covered. Emphasis throughout is given to Lockalloy fabrication and machining techniques and attendant personnel safety precautions. Costs are also examined. Photographs of tested alloy specimens are shown along with the test equipment used.

Duba, R. J.↗
NASA NTRS1979

Strainrange partitioning life predictions of the long time metal properties council creep-fatigue tests

The method of strainrange partitioning is used to predict the cyclic lives of the Metal Properties Council's long time creep-fatigue interspersion tests of several steel alloys. Comparisons are made with predictions based upon the time- and cycle-fraction approach. The method of strainrange partitioning is shown to give consistently more accurate predictions of cyclic life than is given by the time- and cycle-fraction approach.

Saltsman, J. F.↗