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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 433 records · Page 24
All sourcesNASA NTRSDOE OSTI
NASA NTRS1983

Transverse thermal expansion of carbon fiber/epoxy matrix composites

Thermal expansion coefficients and moduli of elasticity have been determined experimentally for a series of epoxy-matrix composites reinforced with carbon and Kevlar fibers. It is found that in the transverse direction the difference between the properties of the fiber and the matrix is not as pronounced as in the longitudinal direction, where the composite properties are fiber-dominated. Therefore, the pattern of fiber packing tends to affect transverse composite properties. The transverse properties of the composites tested are examined from the standpoint of the concept of homogeneity defined as the variation of packing (or lack thereof) throughout a sample.

Helmer, J. F.↗
NASA NTRS2008

Tensile Creep and Fatigue of Sylramic-iBN Melt-Infiltrated SiC Matrix Composites: Retained Properties, Damage Development, and Failure Mechanisms

An understanding of the elevated temperature tensile creep, fatigue, rupture, and retained properties of ceramic matrix composites (CMC) envisioned for use in gas turbine engine applications are essential for component design and life-prediction. In order to quantify the effect of stress, time, temperature, and oxidation for a state-of-the-art composite system, a wide variety of tensile creep, dwell fatigue, and cyclic fatigue experiments were performed in air at 1204 C for the SiC/SiC CMC system consisting of Sylramic-iBN SiC fibers, BN fiber interphase coating, and slurry-cast melt-infiltrated (MI) SiC-based matrix. Tests were either taken to failure or interrupted. Interrupted tests were then mechanically tested at room temperature to determine the residual properties. The retained properties of most of the composites subjected to tensile creep or fatigue were usually within 20% of the as-produced strength and 10% of the as-produced elastic modulus. It was observed that during creep, residual stresses in the composite are altered to some extent which results in an increased compressive stress in the matrix upon cooling and a subsequent increased stress required to form matrix cracks. Microscopy of polished sections and the fracture surfaces of specimens which failed during stressed-oxidation or after the room-temperature retained property test was performed on some of the specimens in order to quantify the nature and extent of damage accumulation that occurred during the test. It was discovered that the distribution of stress-dependent matrix cracking at 1204 C was similar to the as-produced composites at room temperature; however, matrix crack growth occurred over time and typically did not appear to propagate through thickness except at final failure crack. Failure of the composites was due to either oxidation-induced unbridged crack growth, which dominated the higher stress regime (> 179 MPa) or controlled by degradation of the fibers, probably caused by intrinsic creep-induced flaw growth of the fibers or internal attack of the fibers via Si diffusion through the CVI SiC and/or microcracks at the lower stress regime (< 165 MPa).

Morscher, Greg↗
NASA NTRS1984

Analysis of stress-strain, fracture and ductility behavior of aluminum matrix composites containing discontinuous silicon carbide reinforcement

Mechanical properties and stress-strain behavior for several types of commercially fabricated aluminum matrix composites, containing up to 40 vol % discontinuous SiC whisker, nodule, or particulate reinforcement were evaluated. It was found that the elastic modulus of the composites was isotropic, to be independent of type of reinforcement, and to be controlled solely by the volume percentage of SiC reinforcement present. The yield/tensile strengths and ductility were controlled primarily by the matrix alloy and temper condition. Ductility decreased with increasing reinforcement content, however, the fracture strains observed were higher than those reported in the literature for this type of composite. This increase in fracture strain is attributed to cleaner matrix powder and increased mechanical working during fabrication. Conventional aluminum and titanium structural alloys were compared and have shown that the properties of these low cost, lightweight composites have good potential for application to aerospace structures.

Mcdanels, D. L.↗
NASA NTRS2014

Processing and Material Characterization of Continuous Basalt Fiber Reinforced Ceramic Matrix Composites Using Polymer Derived Ceramics.

The need for high performance vehicles in the aerospace industry requires materials which can withstand high loads and high temperatures. New developments in launch pads and infrastructure must also be made to handle this intense environment with lightweight, reusable, structural materials. By using more functional materials, better performance can be seen in the launch environment, and launch vehicle designs which have not been previously used can be considered. The development of high temperature structural composite materials has been very limited due to the high cost of the materials and the processing needed. Polymer matrix composites can be used for temperatures up to 260C. Ceramics can take much higher temperatures, but they are difficult to produce and form in bulk volumes. Polymer Derived Ceramics (PDCs) begin as a polymer matrix, allowing a shape to be formed and cured and then to be pyrolized in order to obtain a ceramic with the associated thermal and mechanical properties. The use of basalt in structural and high temperature applications has been under development for over 50 years, yet there has been little published research on the incorporation of basalt fibers as a reinforcement in the composites. In this study, continuous basalt fiber reinforced PDCs have been fabricated and tested for the applicability of this composite system as a high temperature structural composite material. The oxyacetylene torch testing and three point bend testing have been performed on test panels and the test results are presented.

Polymer Derived Ceramics↗
NASA NTRS1981

Degradation Mechanisms in Aluminum Matrix Composites: Alumina/Aluminum and Boron/Aluminum

The effects of fabrication and long term thermal exposure (up to 10,000 hours at 590 K) on two types of aluminum matrix composites were examined. An alumina/aluminum composite, was made of continuous alpha Al2O3 fibers in a matrix of commercially pure aluminum alloyed with 2.8% lithium. The mechanical properties of the material, the effect of isothermal exposure, cyclic thermal exposure, and fatigue are presented. Two degradation mechanisms are identified. One was caused by formation of a nonstoichiometric alumina during fabrication, the other by a loss of lithium to a surface reaction during long term thermal exposure. The other composite, boron/aluminum, made of boron fibers in an aluminum matrix, was investigated using five different aluminum alloys for the matrices. The mechanical properties of each material and the effect of isothermal and cyclic thermal exposure are presented. The effects of each alloy constituent on the degradation mechanisms are discussed. The effects of several reactions between alloy constituents and boron fibers on the composite properties are discussed.

Olsen, G. C.↗
NASA NTRS1991

Mechanical behavior of a continuous fiber reinforced aluminum matrix composite subjected to transverse and thermal loading

The transverse properties of an aluminum alloy metal matrix composite reinforced by continuous alumina fibers were investigated. The composite is subjected to both mechanical and cyclic thermal loading. The results of an experimental program indicate that the shakedown concept of structural mechanics provides a means of describing the material behavior. When the loading conditions are within the shakedown region, the material finally responds in an elastic manner after initial plastic response, and for loading conditions outside the shakedown region, the material exhibits a rapid incremental plastic strain accumulation. The failure strain varies by an order of magnitude according to the operating conditions. Hence, for high mechanical and low thermal loading, the failure strains is small; for low mechanical and high thermal loading, the failure strain is large.

Jansson, S.↗
NASA NTRS1992

Fiber shape effects on metal matrix composite behavior

The effects of different fiber shapes on the behavior of a SiC/Ti-15 metal matrix composite is computationally simulated. A three-dimensional finite element model consisting of a group of nine unidirectional fibers is used in the analysis. The model is employed to represent five different fiber shapes: a circle, an ellipse, a kidney, and two different cross shapes. The distribution of microstresses and the composite material properties, such as moduli, coefficients of thermal expansion, and Poisson's ratios, are obtained from the finite element analysis for the various fiber shapes. Comparisons of these results are used to determine the sensitivity of the composite behavior to the different fiber shapes and assess their potential benefits. No clear benefits result from different fiber shapes though there are some increases/decreases in isolated properties.

Brown, H. C.↗
NASA NTRS1993

Graphite fiber textile preform/copper matrix composites

This project has the objective of exploring the use of graphite fiber textile preform/copper matrix composites in spacecraft heat transmitting and radiating components. The preforms are to be fabricated by braiding of tows and when infiltrated with copper will result in a 3-D reinforced, near net shape composite with improved specific properties such as lower density and higher stiffness. It is anticipated that the use of textile technology will result in a more robust preform and consequently better final composite; it is hard to anticipate what performance tradeoffs will result, and these will be explored through testing and characterization.

Filatovs, G. J.↗
NASA NTRS1999

Fabrication and Testing of Ceramic Matrix Composite Rocket Propulsion Components

The National Aeronautics and Space Administration (NASA) is pursuing using ceramic matrix composites (CMC) as primary structural components for advanced rocket engines. This endeavor is due to the requirement of increasing safety by two orders of magnitude and reducing costs from $10,000/lb to $1,000/lb both within ten years. Out year goals are even more aggressive. Safety gains, through using CMCS, will be realized by increasing temperature margins, tolerance for extreme thermal transients, and damping capability of components and systems, by using components with lower weight and thermal conductivity, etc. Gains in cost reduction, through using CMCS, are anticipated by enabling higher performance systems, using lighter weight components and systems, enabling 100 mission reusability without system refurbishment, greatly reducing cooling requirements and erosion rates, selecting safe fabrication processes that are ideally cost competitive with metal processes at low volume production, etc. This philosophy contrasts the previous philosophy of rocket engine development focused largely on achieving the highest performance with metals and ablatives -- cost and safety were not the focal point of the initial design. Rocket engine components currently being pursued, largely C/SiC and SiC/SiC, include blisks or rotors, 10 foot by 8 foot nozzle ramps, gas generators, thrust chambers, and upperstage nozzles. The Simplex Turbopump CMC blisk effort has just successfully completed a 4.5 year development and test program. The other components mentioned are in the design or fabrication stage. Although the temperature limits of the CMC materials are not quantified in a realistic environment yet, CMC materials are projected to be the only way to achieve significant safety risks mitigation and cost reductions simultaneously. We, the end-users, material fabricators, technology facilitators, and government organizations are charged with developing and demonstrating a much safer and a lot less costly Earth-to-Orbit full-scale propulsion system by 2005.

Effinger, Michael↗
NASA NTRS2000

Durability Evaluation of a Thin Film Sensor System With Enhanced Lead Wire Attachments on SiC/SiC Ceramic Matrix Composites

An advanced thin film sensor system instrumented on silicon carbide (SiC) fiber reinforced SiC matrix ceramic matrix composites (SiC/SiC CMCs), was evaluated in a Mach 0.3 burner rig in order to determine its durability to monitor material/component surface temperature in harsh environments. The sensor system included thermocouples in a thin film form (5 microns thick), fine lead wires (75 microns diameter), and the bonds between these wires and the thin films. Other critical components of the overall system were the heavy, swaged lead wire cable (500 microns diameter) that contained the fine lead wires and was connected to the temperature readout, and ceramic attachments which were bonded onto the CMCs for the purpose of securing the lead wire cables, The newly developed ceramic attachment features a combination of hoops made of monolithic SiC or SiC/SiC CMC (which are joined to the test article) and high temperature ceramic cement. Two instrumented CMC panels were tested in a burner rig for a total of 40 cycles to 1150 C (2100 F). A cycle consisted of rapid heating to 1150 C (2100 F), a 5 minute hold at 1150 C (2100 F), and then cooling down to room temperature in 2 minutes. The thin film sensor systems provided repeatable temperature measurements for a maximum of 25 thermal cycles. Two of the monolithic SiC hoops debonded during the sensor fabrication process and two of the SiC/SiC CMC hoops failed during testing. The hoops filled with ceramic cement, however, showed no sign of detachment after 40 thermal cycle test. The primary failure mechanism of this sensor system was the loss of the fine lead wire-to-thin film connection, which either due to detachment of the fine lead wires from the thin film thermocouples or breakage of the fine wire.

Lei, Jih-Fen↗
NASA NTRS2001

Thermal Conductivity and Thermal Gradient Cyclic Behavior of Refractory Silicate Coatings on SiC/SiC Ceramic Matrix Composites

Plasma-sprayed mullite and BSAS coatings have been developed to protect SiC/SiC ceramic matrix composites from high temperature environmental attack. In this study, thermal conductivity and thermal barrier functions of these coating systems are evaluated using a laser high-heat-flux test rig. The effects of water vapor on coating thermal conductivity and durability are studied by using alternating furnace and laser thermal gradient cyclic tests. The influence of laser high thermal-gradient cycling on coating failure modes is also investigated.

Zhu, Dongming↗
NASA NTRS2002

Characterization of Ceramic Matrix Composite Fasteners Exposed in a Combustion Liner Rig Test

Combustion tests on SiC/SiC CMC (Ceramic Matrix Composite) components were performed in an aircraft combustion environment using the Rich-burn, Quick-quench, Lean-burn (RQL) sector rig. SiC/SiC fasteners were used to attach several of these components to the metallic rig structure. The effect of combustion exposure on the fastener material was characterized via microstructural examination. Fasteners were also destructively tested, after combustion exposure, and the failure loads of fasteners exposed in the sector rig were compared to those of as-manufactured fasteners. Combustion exposure reduced the fastener failure load by 50% relative to the as-manufactured fasteners for exposure times ranging from 50 to 260 hours. The fasteners exposed in the combustion environment demonstrated failure loads that varied with failure mode. Fasteners that had the highest average failure load, failed in the same manner as the unexposed fasteners.

Verrilli, Michael J.↗
NASA NTRS2003

Measurements of Erosion Wear Volume Loss on Bare and Coated Polymer Matrix Composites

An investigation was conducted to examine the erosion behavior of uncoated and coated polymer matrix composite (PMC) specimens subjected to solid particle impingement using air jets. The PMCs were carbon-Kevlar (DuPont, Wilmington, DE) fiber-epoxy resin composites with a temperature capability up to 393 K (248 F). Tungsten carbide-cobalt (WC-Co) was the primary topcoat constituent. Bondcoats were applied to the PMC substrates to improve coating adhesion; then, erosion testing was performed at the University of Cincinnati. All erosion tests were conducted with Arizona road-dust (ARD), impinging at angles of 20 and 90 on both uncoated and two-layer coated PMCs at a velocity of 229 m/s and at a temperature of 366 K (200 F). ARD contains primarily 10-m aluminum oxide powders. Vertically scanning interference microscopy (noncontact, optical profilometry) was used to evaluate surface characteristics, such as erosion wear volume loss and depth, surface topography, and surface roughness. The results indicate that noncontact, optical interferometry can be used to make an accurate determination of the erosion wear volume loss of PMCs with multilayered structures while preserving the specimens. The two-layered (WC-Co topcoat and metal bondcoat) coatings on PMCs remarkably reduced the erosion volume loss by a factor of approximately 10. The tenfold increase in erosion resistance will contribute to longer PMC component lives, lower air friction, reduced related breakdowns, decreased maintenance costs, and increased PMC reliability. The decrease in the surface roughness of the coated vanes will lead to lower air friction and will subsequently reduce energy consumption. Eventually, the coatings could lead to overall economic savings.

Miyoshi, Kazuhisa↗
NASA NTRS1990

Ceramic Matrix Composites (CMC) Life Prediction Development

Advanced launch systems will very likely incorporate fiber reinforced ceramic matrix composites (CMC) in critical propulsion and airframe components. The use of CMC will save weight, increase operating margin, safety and performance, and improve reuse capability. For reusable and single mission use, accurate life prediction is critical to success. The tools to accomplish this are immature and not oriented toward the behavior of carbon fiber reinforced silicon carbide (C/SiC), the primary system of interest for many applications. This paper describes an approach and progress made to satisfy the need to develop an integrated life prediction system that addresses mechanical durability and environmental degradation.

Levine, Stanley R.↗
NASA NTRS1992

Effects of fiber and interfacial layer architectures on the thermoplastic response of metal matrix composites

Examined here is the effect of fiber and interfacial layer morphologies on thermal fields in metal matrix composites (MMCs). A micromechanics model based on an arbitrarily layered concentric cylinder configuration is used to calculate thermal stress fields in MMCs subjected to spatially uniform temperature changes. The fiber is modelled as a layered material with isotropic or orthotropic elastic layers, whereas the surrounding matrix, including interfacial layers, is treated as a strain-hardening, elastoplastic, von Mises solid with temperature-dependent parameters. The solution to the boundary-value problem of an arbitrarily layered concentric cylinder under the prescribed thermal loading is obtained using the local/global stiffness matrix formulation originally developed for stress analysis of multilayered elastic media. Examples are provided that illustrate how the morphology of the SCS6 silicon carbide fiber and the use of multiple compliant layers at the fiber/matrix interface affect the evolution of residual stresses in SiC/Ti composites during fabrication cool-down.

Pindera, Marek-Jerzy↗
NASA NTRS2003

Ultrasonic Assessment of Impact-Induced Damage and Microcracking in Polymer Matrix Composites

The main objective of this NASA Faculty Awards for Research (FAR) project is to conduct ultrasonic assessment of impact-induced damage and microcracking in fiber-metal laminated (FML) composites at various temperatures. It is believed that the proposed study of impact damage assessment on FML composites will benefit several NASA's missions and current interests, such as ballistic impact testing of composite fan containment and high strain rate deformation modeling of polymer matrix composites. Impact-induced damage mechanisms in GLARE and ARALL fiber-metal laminates subject to instrumented drop-weight impacts at various temperatures were studied. GLARE and ARALL are hybrid composites made of alternating layers of aluminum and glass- (for GLARE) and aramid- (for ARALL) fiber reinforced epoxy. Damage in pure aluminum panels impacted by foreign objects was mainly characterized by large plastic deformation surrounding a deep penetration dent. On the other hand, plastic deformation in fiber-metal laminates was often not as severe although the penetration dent was still produced. The more stiff fiber-reinforced epoxy layers provided better bending rigidity; thus, enhancing impact damage tolerance. Severe cracking, however, occurred due to the use of these more brittle fiber-reinforced epoxy layers. Fracture patterns, e.g., crack length and delamination size, were greatly affected by the lay-up configuration rather than by the number of layers, which implies that thickness effect was not significant for the panels tested in this study. Immersion ultrasound techniques were then used to assess damages generated by instrumented drop-weight impacts onto these fiber-metal laminate panels as well as 2024-T3 aluminum/cast acrylic sandwich plates adhered by epoxy. Depending on several parameters, such as impact velocity, mass, temperature, laminate configuration, sandwich construction, etc., various types of impact damage were observed, including plastic deformation, radiating cracks emanating from the impact site, ring cracks surrounding the impact site, partial and full delamination, and combinations of these damages.

Gyekanyesi, John↗
NASA NTRS2007

Double Vacuum Bag Process for Resin Matrix Composite Manufacturing

A double vacuum bag molding assembly with improved void management and laminate net shape control which provides a double vacuum enviromnent for use in fabricating composites from prepregs containing air and/or volatiles such as reactive resin matrix composites or composites from solvent containing prepregs with non-reactive resins matrices. By using two vacuum environments during the curing process, a vacuum can be drawn during a B-stage of a two-step cycle without placing the composite under significant relative pressure. During the final cure stage, a significant pressure can be applied by releasing the vacuum in one of the two environments. Inner and outer bags are useful for creating the two vacuum environments with a perforated tool intermediate the two. The composite is placed intermediate a tool plate and a caul plate in the first environment with the inner bag and tool plate defining the first environment. The second environment is characterized by the outer bag which is placed over the inner bag and the tool plate.

Hou, Tan-Hung↗