Search NASA⌕ Search

SEARCH · Search NASA

Results for “REINFORCING FIBER”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 523 records · Page 29
All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

Novel Approach for Positioning Sensor Lead Wires on SiC-Based Monolithic Ceramic and FRCMC Components/Subcomponents Having Flat and Curved Surfaces

A novel attachment approach for positioning sensor lead wires on silicon carbide-based monolithic ceramic and fiber reinforced ceramic matrix composite (FRCMC) components has been developed. This approach is based on an affordable, robust ceramic joining technology, named ARCJoinT, which was developed for the joining of silicon carbide-based ceramic and fiber reinforced composites. The ARCJoinT technique has previously been shown to produce joints with tailorable thickness and good high temperature strength. In this study, silicon carbide-based ceramic and FRCMC attachments of different shapes and sizes were joined onto silicon carbide fiber reinforced silicon carbide matrix (SiC/ SiC) composites having flat and curved surfaces. Based on results obtained in previous joining studies. the joined attachments should maintain their mechanical strength and integrity at temperatures up to 1350 C in air. Therefore they can be used to position and secure sensor lead wires on SiC/SiC components that are being tested in programs that are focused on developing FRCMCs for a number of demanding high temperature applications in aerospace and ground-based systems. This approach, which is suitable for installing attachments on large and complex shaped monolithic ceramic and composite components, should enhance the durability of minimally intrusive high temperature sensor systems. The technology could also be used to reinstall attachments on ceramic components that were damaged in service.

Kiser, J. Douglas↗
NASA NTRS2000

Raman Study of Uncoated and P-bn/sic-coated Hi-nicalon Reinforced Celsian Matrix Composites: Residual Stress in the Fibers - Part 2

Band shifts on Raman spectra were used to assess, at a microscopic scale, the residual strain existing in Hi-Nicalon fibers reinforcing celsian matrix composites. Uncoated as well as p-BN/SiC- and p-B(Si)N/SiC-coated Hi-Nicalon fibers were used as the reinforcements. We unambiguously conclude that the fibers are in a state of compressive residual stress. Quantitative determination of the residual stress was made possible by taking into account the heating induced by laser probing and by using a reference line, of fixed wavenumber. We found fiber compressive residual stress values between 110 and 960 MPa depending on the fiber/matrix coating in the composite. A stress relaxation-like phenomenon was observed at the surface of p-BN/SiC-coated Hi-Nicalon fibers whereas the uncoated or p-B(Si)N/SiC-coated Hi-Nicalon fibers did not show any stress relaxation in the Celsian matrix composites.

Gouadec, Gwenael↗
NASA NTRS1998

Effects of Fiber/Matrix Interface and its Composition on Mechanical Properties of Hi Nicalon/Celsian Composites

Fiber-reinforced ceramic matrix composites (CMC) are prospective candidate materials for high temperature structural applications in aerospace, energy conservation, power generation, nuclear, petrochemical, and other industries. At NASA Lewis, we are investigating celsian matrix composites reinforced with various types of silicon carbide fibers. The objective of the present study was to investigate the effects of fiber/matrix interface and its composition on the mechanical properties of silicon carbide (Hi-Nicalon) fiber-reinforced celsian matrix composites.

Bansal, Narottam P.↗
NASA NTRS2000

Use of Guided Acoustic Waves to Assess the Effects of Thermal-Mechanical Cycling on Composite Stiffness

The introduction of new, advanced composite materials into aviation systems requires it thorough understanding of the long-term effects of combined thermal and mechanical loading. As part of a study to evaluate the effects of thermal-mechanical cycling, it guided acoustic (Lamb) wave measurement system was used to measure the bending and out-of-plane stiffness coefficients of composite laminates undergoing thermal-mechanical loading. The system uses a pulse/receive technique that excites an antisymmetric Lamb mode and measures the time-of-flight over a wide frequency range. Given the material density and plate thickness, the bending and out-of-plane shear stiffnesses are calculated from a reconstruction of the velocity dispersion curve. A series of 16 and 32-ply composite laminates were subjected to it thermal-mechanical loading profile in load frames equipped with special environmental chambers. The composite systems studied were it graphite fiber reinforced amorphous thermoplastic polyimide and it graphite fiber reinforced bismaleimide thermoset. The samples were exposed to both high and low temperature extremes its well as high and low strain profiles. The bending and out-of-plane stiffnesses for composite sample that have undergone over 6,000 cycles of thermal-mechanical loading are reported. The Lamb wave generated elastic stiffness results have shown decreases of up to 20% at 4,936 loading cycles for the graphite/thermoplastic samples and up to 64% at 4,706 loading cycles for the graphite/thermoset samples.

Seale, Michael D.↗
NASA NTRS1992

Characteristics of laminates with delamination control strips

Tough resin is needed to resist delamination crack propagation. However, modulus often has to be compromised because it is difficult to retain both high modulus and toughness in a matrix material. A potential solution is to use a hybrid system in which tough resin strips are included within a conventional matrix composite. By adjusting the spacing of the tough resin strips, maximum delamination size can be controlled. Experimental results for impact damage and subsequent damage propagation in laminates containing tough resin strips are reported. Plain adhesive strips and fiber-reinforced tough resin composite strips were used in constructing the hybrid laminates. Test results indicated that size of delamination inflicted by impact was confined between the tough resin strips. As a result, significantly increased residual compressive strength was obtained. Impacted laminates containing tough resin strips were also fatigue tested. It was found that these strips reduced the growth of the impact damage area relative to the growth seen in coupons with no tough resin strips. Damage growth from an open hole under tension fatigue was evaluated using both tough resin strips and glass fiber reinforced tough resin strips. Unreinforced tough resin strips retarded delamination growth from the open hole, but did not stop matrix cracks growing in the fiber direction. Fiber reinforced tough resin strips did not contain axial delamination growth from the open hole. However, they did act as crack arresters, stopping the through-the-thickness tension crack originating from the hole.

Sun, C. T.↗
NASA NTRS2010

Economical Fabrication of Thick-Section Ceramic Matrix Composites

A method was developed for producing thick-section [>2 in. (approx.5 cm)], continuous fiber-reinforced ceramic matrix composites (CMCs). Ultramet-modified fiber interface coating and melt infiltration processing, developed previously for thin-section components, were used for the fabrication of CMCs that were an order of magnitude greater in thickness [up to 2.5 in. (approx.6.4 cm)]. Melt processing first involves infiltration of a fiber preform with the desired interface coating, and then with carbon to partially densify the preform. A molten refractory metal is then infiltrated and reacts with the excess carbon to form the carbide matrix without damaging the fiber reinforcement. Infiltration occurs from the inside out as the molten metal fills virtually all the available void space. Densification to <5 vol% porosity is a one-step process requiring no intermediate machining steps. The melt infiltration method requires no external pressure. This prevents over-infiltration of the outer surface plies, which can lead to excessive residual porosity in the center of the part. However, processing of thick-section components required modification of the conventional process conditions, and the means by which the large amount of molten metal is introduced into the fiber preform. Modification of the low-temperature, ultraviolet-enhanced chemical vapor deposition process used to apply interface coatings to the fiber preform was also required to accommodate the high preform thickness. The thick-section CMC processing developed in this work proved to be invaluable for component development, fabrication, and testing in two complementary efforts. In a project for the Army, involving SiC/SiC blisk development, nominally 0.8 in. thick x 8 in. diameter (approx. 2 cm thick x 20 cm diameter) components were successfully infiltrated. Blisk hubs were machined using diamond-embedded cutting tools and successfully spin-tested. Good ply uniformity and extremely low residual porosity (<2 percent) were achieved, the latter being far lower than that achieved with SiC matrix composites fabricated via CVI or PIP. The pyrolytic carbon/zirconium nitride interface coating optimized in this work for use on carbon fibers was incorporated in the SiC/SiC composites and yielded a >41 ksi (approx. 283 MPa) flexural strength.

Babcock, Jason↗
NASA NTRS2005

SiC/SiC Ceramic Matrix Composites Developed for High-Temperature Space Transportation Applications

Researchers at the NASA Glenn Research Center have been developing durable, high-temperature ceramic matrix composites (CMCs) with silicon carbide (SiC) matrices and SiC or carbon fibers for use in advanced reusable launch vehicle propulsion and airframe applications in the Next Generation Launch Technology (NGLT) Program. These CMCs weigh less and are more durable than competing metallic alloys, and they are tougher than silicon-based monolithic ceramics. Because of their high specific strength and durability at high temperatures, CMCs such as C/SiC (carbon- fiber-reinforced silicon carbide) and SiC/SiC (silicon-carbide-fiber-reinforced silicon carbide) may increase vehicle performance and safety significantly and reduce the cost of transporting payloads to orbit.

Kiser, J. Douglas↗
NASA NTRS2000

Affordable, Robust Ceramic Joining Technology (ARCJoinT) Given 1999 R and D 100 Award

Advanced ceramics and fiber-reinforced ceramic matrix composites with high strength and toughness, good thermal conductivity, thermal shock resistance, and oxidation resistance are needed for high-temperature structural applications in advanced high-efficiency and high-performance engines, space propulsion components, and land-based systems. The engineering designs of these systems require the manufacturing of large parts with complex shapes, which are either quite expensive or impossible to fabricate. In many instances, it is more economical to build complex shapes by joining together simple geometrical shapes. Thus, joining has been recognized as an enabling technology for the successful utilization of advanced ceramics and fiber-reinforced composite components in high-temperature applications. However, such joints must retain their structural integrity at high temperatures and must have mechanical strength and environmental stability comparable to those of the bulk materials. In addition, the joining technique should be robust, practical, and reliable. ARCJoinT, which is based on the reaction-forming approach, is unique in terms of producing joints with tailorable microstructures. The formation of joints by this approach is attractive since the thermomechanical properties of the joint interlayer can be tailored to be very close to those of the base materials. In addition, high-temperature fixturing is not needed to hold the parts at the infiltration temperature. The joining process begins with the application of a carbonaceous mixture in the joint area, holding the items to be joined in a fixture, and curing at 110 to 120 C for 10 to 20 min. This step fastens the pieces together. Then, silicon or a silicon alloy in tape, paste, or slurry form is applied around the joint region and heated to 1250 to 1425 C (depending on the type of infiltrant) for 10 to 15 min. The molten silicon or silicon-refractory metal alloy reacts with carbon to form silicon carbide with controllable amounts of silicon and other phases as determined by the alloy composition. Joint thickness can be readily controlled through adjustments of the properties of the carbonaceous paste and the applied fixturing force. The photograph shows various shapes of silicon-carbide-based ceramics and fiberreinforced composites that have been joined using ARCJoinT. Thermomechanical and thermochemical characterization of joints is underway for a wide variety of silicon-carbidebased advanced ceramics and fiber-reinforced composites under the hostile environments that will be encountered in engine applications. ARCJoinT, which was developed by researchers at the NASA Glenn Research Center at Lewis Field, received R&D Magazine's prestigious R&D 100 Award in 1999.

Singh, Mrityunjay↗
NASA NTRS1989

Techniques for measurement of the thermal expansion of advanced composite materials

Techniques available to measure small thermal displacements in flat laminates and structural tubular elements of advanced composite materials are described. Emphasis is placed on laser interferometry and the laser interferometric dilatometer system used at the National Aeronautics and Space Administration (NASA) Langley Research Center. Thermal expansion data are presented for graphite-fiber reinforced 6061 and 2024 aluminum laminates and for graphite fiber reinforced AZ91 C and QH21 A magnesium laminates before and after processing to minimize or eliminate thermal strain hysteresis. Data are also presented on the effects of reinforcement volume content on thermal expansion of silicon-carbide whisker and particulate reinforced aluminum.

Tompkins, Stephen S.↗
NASA NTRS1991

Experiment M0003-10: Advanced composites

The Advanced composites Sub-experiment includes numerous metal matrix composites, primarily graphite fiber reinforced aluminum and magnesium, and several classes of graphite fiber reinforced organic matrix (epoxy, polysulfone, polyimide) composites with and/or without thermal control or protective coatings. The experiment included over 500 flight samples, mounted on the leading and trailing edges of LDEF, and a full complement of the laboratory control samples. Preliminary results are presented reviewing changes in the visual appearance of the samples, atomic oxygen erosion of the organic matrix composites, contamination effects, and the effects of micrometeorite impacts and thermal fatigue on the metal matrix composites.

Steckel, Gary L.↗
NASA NTRS1991

Advanced materials for space nuclear power systems

The overall philosophy of the research was to develop and characterize new high temperature power conversion and radiator materials and to provide spacecraft designers with material selection options and design information. Research on three candidate materials (carbide strengthened niobium alloy PWC-11 for fuel cladding, graphite fiber reinforced copper matrix composites for heat rejection fins, and tungsten fiber reinforced niobium matrix composites for fuel containment and structural supports considered for space power system applications is discussed. Each of these types of materials offers unique advantages for space power applications.

Titran, Robert H.↗
NASA NTRS1991

Advanced materials for space nuclear power systems

The overall philosophy of the research was to develop and characterize new high temperature power conversion and radiator materials and to provide spacecraft designers with material selection options and design information. Research on three candidate materials (carbide strengthened niobium alloy PWC-11 for fuel cladding, graphite fiber reinforced copper matrix composites for heat rejection fins, and tungsten fiber reinforced niobium matrix composites for fuel containment and structural supports) considered for space power system applications is discussed. Each of these types of materials offers unique advantages for space power applications.

Titran, Robert H.↗
NASA NTRS1998

High Temperature Mechanical Characterization of Ceramic Matrix Composites

A high temperature mechanical characterization laboratory has been assembled at NASA Lewis Research Center. One contribution of this work is to test ceramic matrix composite specimens in tension in environmental extremes. Two high temperature tensile testing systems were assembled. The systems were assembled based on the performance and experience of other laboratories and meeting projected service conditions for the materials in question. The systems use frames with an electric actuator and a center screw. A PC based data acquisition and analysis system is used to collect and analyze the data. Mechanical extensometers are used to measure specimen strain. Thermocouples, placed near the specimen, are used to measure the specimen gage section temperature. The system for testing in air has a resistance element furnace with molybdenum disilicide elements and pneumatic grips with water cooling attached to hydraulic alignment devices. The system for testing in an inert gas has a graphite resistance element furnace in a chamber with rigidly mounted, water cooled, hydraulically actuated grips. Unidirectional SiC fiber reinforced reaction bonded Si3N4 and triaxially woven, two dimensional, SiC fiber reinforced enhanced SiC composites were tested in unidirectional tension. Theories for predicting the Young's modulus, modulus near the ultimate strength, first matrix cracking stress, and ultimate strength were applied and evaluated for suitability in predicting the mechanical behavior of SiC/RBSN and enhanced SiC/SiC composites. The SiC/RBSN composite exhibited pseudo tough behavior (increased area under the stress/strain curve) from 22 C to 1500 C. The rule of mixtures provides a good estimate of the Young's modulus of the SiC/RBSN composite using the constituent properties from room temperature to 1440 C for short term static tensile tests in air or nitrogen. The rule of mixtures significantly overestimates the secondary modulus near the ultimate strength. The ACK theory provides the best approximation of the first matrix cracking stress when residual stresses are ignored. The theory of Cao and Thouless, based on Weibull statistics, gave the best prediction for the composite ultimate strength. The enhanced SiC/SiC composite exhibited nonlinear stress/strain behavior from 24 C to 1370 C in air with increased ultimate strain when compared to monolithic SiC. The theory of Yang and Chou with the assumption of a frictional fiber/matrix interface provided the best estimate of the Young's modulus. The theory of Cao and Thouless gave the best estimate for the ultimate strength.

Gyekenyesi, John Z.↗
NASA NTRS2005

Improvements in Fabrication of Sand/Binder Cores for Casting

Three improvements have been devised for the cold-box process, which is a special molding process used to make sand/binder cores for casting hollow metal parts. These improvements are: The use of fiber-reinforced composite binder materials (in contradistinction to the non-fiber-reinforced binders used heretofore), The substitution of a directed-vortex core-blowing subprocess for a prior core-blowing process that involved a movable gassing plate, and The use of filters made from filtration-grade fabrics to prevent clogging of vents. For reasons that exceed the scope of this article, most foundries have adopted the cold-box process for making cores for casting metals. However, this process is not widely known outside the metal-casting industry; therefore, a description of pertinent aspects of the cold-box process is prerequisite to a meaningful description of the aforementioned improvements. In the cold-box process as practiced heretofore, sand is first mixed with a phenolic resin (considered to be part 1 of a three-part binder) and an isocyanate resin (part 2 of the binder). Then by use of compressed air, the mixture is blown into a core box, which is a mold for forming the core. Next, an amine gas (part 3 of the binder) that acts as a catalyst for polymerization of parts 1 and 2 is blown through the core box. Alternatively, a liquid amine that vaporizes during polymerization can be incorporated into the sand/resin mixture. Once polymerization is complete, the amine gas is purged from the core box by use of compressed air. The finished core is then removed from the core box.

Bakhitiyarov, Sayavur I.↗
NASA NTRS1994

Post-impact fatigue of cross-plied, through-the-thickness reinforced carbon/epoxy composites

An experimental investigation of the post-impact fatigue response of integrally woven carbon/epoxy composites was conducted. Five different through-the-thickness (TTT) reinforcing fibers were used in an experimental textile process that produced an integrally woven (0/90/0/90/0/90/0/90/0)(sub T) ply layup with 21K AS4 carbon tow fiber. The resin was Hercules 3501-6, and the five TTT reinforcing fibers were Kevlar, Toray carbon, AS4 carbon, glass, and IM6 carbon. The purpose of this investigation was to study the post-impact fatigue response of these material systems and to identify the optimum TTT fiber. Samples were impacted with one half inch diameter aluminum balls with an average velocity of 543 ft/sec. Post-impact static compression and constant amplitude tension-compression fatigue tests were conducted. Fatigue tests were conducted with a loading ratio of R=-5, and frequency of 4 Hz. Damage growth was monitored using x-radiographic and sectioning techniques and by examining the stress-strain response (across the impact site) throughout the fatigue tests. The static compressive stress versus far-field strain response was nearly linear for all material groups. All the samples had a transverse shear failure mode. The average compressive modulus (from far-field strain) was about 10 Msi. The average post-impact static compressive strength was about 35.5 Ksi. The IM6 carbon sample had a strength of over 40 Ksi, more than 16 percent stronger than average. There was considerable scatter in the S-N data. However, the IM6 carbon samples clearly had the best fatigue response. The response of the other materials, while worse than IM6 carbon, could not be ranked definitively. The initial damage zones caused by the impact loading and damage growth from fatigue loading were similar for all five TTT reinforcing materials. The initial damage zones were circular and consisted of delaminations, matrix cracks and ply cracks. Post-impact fatigue loading caused delamination growth, ply cracking and fiber bundle failures, typically 45 deg from impact load direction. During the initial 97 percent of fatigue life, delaminations, ply cracks and fiber bundle failures primarily grew at and near the impact site. During the final 3 percent of life, damage grew rapidly transverse to the loading direction as a through-the-thickness transverse shear failure. The stress-strain response was typically linear during the initial 50 percent of life, and stiffness dropped about 20 percent during this period. During the next 47 percent of life, stiffness dropped about 34 percent, and the stress-strain response was no longer linear. The stiffness decreased about 23 percent during the final 3 percent of life. These trends were typical of all the materials tested. Therefore, by monitoring stiffness loss, fatigue failure could be accurately anticipated.

Serdinak, Thomas E.↗
NASA NTRS2002

Effect of Environment on the Stress- Rupture Behavior of a C/SiC Composite Studied

Advanced reusable launch vehicles will likely incorporate fiber-reinforced ceramic matrix composites (CMC's) in critical propulsion and airframe components. The use of CMC's is highly desirable to save weight, improve reuse capability, and increase performance. One of the candidate CMC materials is carbon-fiber-reinforced silicon carbide (C/SiC). In potential propulsion applications, such as turbopump rotors and nozzle exit ramps, C/SiC components will be subjected to a service cycle that includes mechanical loading under complex, high-pressure environments containing hydrogen, oxygen, and steam. Degradation of both the C fibers and the SiC matrix are possible in these environments. The objective of this effort was to evaluate the mechanical behavior of C/SiC in various environments relevant to reusable launch vehicle applications. Stress-rupture testing was conducted at the NASA Glenn Research Center on C/SiC specimens in air and steam-containing environments. Also, the oxidation kinetics of the carbon fibers that reinforce the composite were monitored by thermogravimetric analysis in the same environments and temperatures used for the stress-rupture tests of the C/SiC composite specimens. The stress-rupture lives obtained for C/SiC tested in air and in steam/argon mixtures are shown in the following bar chart. As is typical for most materials, lives obtained at the lower temperature (600 C) are longer than for the higher temperature (1200 C). The effect of environment was most pronounced at the lower temperature, where the average test duration in steam at 600 C was at least 30 times longer than the lives obtained in air. The 1200 C data revealed little difference between the lives of specimens tested in air and steam at atmospheric pressure.

Verrilli, Michael J.↗