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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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1978

A temperature dependent fatigue failure criterion for graphite/epoxy laminates

A fatigue failure criterion applicable to composite materials is developed and applied to predict the fatigue behavior of graphite/epoxy laminates with particular emphasis on the influence of temperature. Tensile stress-strain curves and tension-tension fatigue curves for various unidirectional, angle-ply and symmetrically balanced laminates were developed at test temperatures of 25 C, 74 C, and 114 C. For most laminates a reduction in both static strength and fatigue strength is observed with increasing temperature. This reduction appeared more severe in fatigue loading than in static tensile loading and most severe where the shear stress in the lamina is the dominate failure mode. Through an analytical formulation of shifting functions for the influences of temperature, all fatigue data are shown to be capable of being reduced to a single reference curve at some temperature. Examples are given which demonstrate the capability of the fatigue failure criterion to predict failure of complex symmetrically balanced laminates from relevant parameters obtained from the observed behavior of unidirectional and angle-ply laminates.

Rotem, A.↗
NASA NTRS1979

A temperature dependent fatigue failure criterion for graphite/epoxy laminates

A fatigue failure criterion applicable to composite materials is developed and applied to predict the fatigue behavior of graphite/epoxy laminates with particular emphasis on the influence of temperature. Tensile stress-strain curves and tension-tension fatigue curves for various unidirectional, angle-ply and symmetrically balanced laminates were developed at test temperatures of 25 C, 74 C and 114 C. In general for most laminates a reduction in both static strength and fatigue strength is observed with increasing temperature. This reduction appeared more severe in fatigue loading than in static tensile loading and most severe where the shear stress in the lamina is the dominant failure mode. Through an analytical formulation of shifting functions for the influences of temperature, all fatigue data are shown to be capable of being reduced to a single reference curve at some temperature. Additionally, examples are given which demonstrate the capability of the fatigue failure criterion to predict failure of complex symmetrically balanced laminates from relevant parameters obtained from the observed behavior of unidirectional and angle-ply laminates.

Rotem, A.↗
NASA NTRS1982

Fatigue behavior of SiC reinforced Ti/6Al-4V/ at 650 C

Axial, low cycle fatigue properties of 25 and 44 fiber vol pct SiC/Ti(6Al-4V) composites, measured at 650 C, were compared with the fatigue properties of unreinforced Ti(6Al-4V) at the same temperature. A prior study of the fatigue behavior of this composite system at room temperature indicated that the SiC fiber reinforcement did not provide the anticipated improvement of fatigue resistance of this alloy. At 650 C, the composite fatigue properties degraded somewhat from those at room temperature. However, these properties degraded more for the unreinforced matrix at 650 C with the result that the composite fatigue strength was two to three times the fatigue strength of the matrix alloy. The reasons for this reversal are discussed in terms of crack initiation at broken fibers and residual matrix stresses.

Bhatt, R. T.↗
NASA NTRS1984

Residual strength of five boron/aluminum laminates with crack-like notches after fatigue loading

Boron/aluminum specimens were made with crack-like slits in the center and with various proportions of 0 and + or - 45 deg plies. They were fatigue loaded and then fractured to determine their residual strengths. The fatigue loads were generally in the range of 60 to 80 percent of the static tensile strength of the specimen as determined from a previous study, and the stress ratio was .05. For virtually all of the specimens the fatigue loading was continued for 100,000 cycles. The specimens were radiographed after the fatigue loading to determine the nature of the fatigue damage. A few specimens were sectioned and examined in a scanning electron microscope after being radiographed in order to verify the interpretation of the radiographs and also to get a better insight into the nature of the fatigue damage. The results indicate that the fatiguing does not significantly affect the strength of the specimens tested. The results of the radiography and of the scanning electron microscopy indicate that the 45 deg plies suffer extensive damage in the form of split and broken fibers and matrix cracking in the vicinity of the ends of the split. By contrast, the only significant damage to the 0 deg plies was a single 0 deg matric crack growing from the ends of the slit and between the 0 deg fibers.

Simonds, R. A.↗
NASA NTRS1984

Fatigue behavior of continuous fiber silicon-carbide-aluminum composites

Four lay-ups of continuous fiber silicon carbide (SCS2) fiber/aluminum matrix composites were tested to assess fatigue mechanisms including stiffness loss when cycled below their respective fatigue limits. The lay-ups were 0 (sub 8), 0(sub 2)/ + or - 45 (sub 2s), 0/90 (sub 2s),and 0/ + or 45/90 (subs). The data were compared with predictions from the author's previously published shakedown model which predicts fatigue-induced stiffness loss in metal matrix composites. A fifth lay-up, + or - 45 (sub 2s), was tested to compare shakedown and fatigue limits. The particular batch of silicon-carbide fibers tested in this program had a somewhat lower modulus (340 GPa) than expected and displayed poor bonding to the aluminum matrix. Good agreement was obtained between the stiffness loss model and the test data. The fatigue damage below the fatigue limit was primarily in the form of matrix cracking. The fatigue limit corresponded to the laminate shakedown for the + or - 45 (sub 2s) laminate.

Johnson, W. S.↗
NASA NTRS1985

Fatigue life extension

Potential fatigue rejuvenation processes were carried out on fatigue-damaged material both with and without observable surface-connected fatigue cracks. The fatigue life of fatigue-damaged MAR-M246(Hf)(DS), a directionally solidified nickel-base superalloy used in turbine airfoils, was extended by reheat treatment. The fatigue life of fatigue-cracked Inconel 718, a wrought nickel-base superalloy used in a wide variety of advanced rocket engine components, was extended by electron-beam welding to close off the surface-connected crack, followed by hot isostatic pressing and reheat treatment.

Matejczyk, D. E.↗
NASA NTRS1985

Creep-fatigue behavior of NiCoCrAlY coated PWA 1480 superalloy single crystals

Single crystal specimens of a Ni base superalloy, PWA 1480, with a low pressure plasma sprayed NiCoCrAlY coating were tested in various 0.1 Hz fatigue and creep fatigue cycles both at 1015 and 1050 C. Creep fatigue tests of the cp, pc, and cc types were conducted with various constant total strain ranges employing creep dwells at various constant stresses. Considerable cyclic softening occurred as was evidenced particularly by rapidly increasing creep rates in the creep fatigue tests. The cycle time in the creep fatigue tests typically decreased by more than 80 percent at 0.5 N sub f. Though cyclic life did correlate with delta epsilon sub in a better correlation existed with sub f for both the fatigue and creep fatigue tests, and poor correlations were observed with either sigma sub max or the average cycle time. A model containing both delta sigma and delta sigma (sub in), N sub f = alpha delta sigma (sub in) beta delta sigma gamma, with best fit values of sigma for each cycle type, but the same values of beta and gamam, was found to provide good correlations. Life lines were not greatly different among the cycle types, differing only by a factor of about three. The cp cycle life line was lowest for both test temperatures, however among the other three cycle types there was no consistent ranking. For all test types failure occurred predominately by multiple internal cracking originating at pores. The strong correlation of life with delta sigma may reflect a significant crack growth period in the life of the specimens.

Miner, R. V.↗
NASA NTRS1985

Elevated temperature biaxial fatigue

A 3 year experimental program for studying elevated temperature biaxial fatigue of a nickel based alloy Hastelloy-X has been completed. A new high temperature fatigue test facility with unique capabilities has been developed. Effort was directed toward understanding multiaxial fatigue and correlating the experimental data to the existing theories of fatigue failure. The difficult task of predicting fatigue lives for nonproportional loading was used as an ultimate test for various life prediction methods being considered. The primary means of reaching improved understanding were through several critical nonproportional loading experiments. The direction of cracking observed on failed specimens was also recorded and used to guide the development of the theory. Cyclic deformation responses were permanently recorded digitally during each test. It was discovered that the cracking mode switched from primarily cracking on the maximum shear planes at room temperature to cracking on the maximum normal strain planes at 649 C. In contrast to some other metals, loading path in nonproportional loading had little effect on fatigue lives. Strain rate had a small effect on fatigue lives at 649 C. Of the various correlating parameters the modified plastic work and octahedral shear stress were the most successful.

Jordan, E. H.↗
NASA NTRS1986

Fatigue behavior of continuous-fiber silicon carbide/aluminum composites

Four lay-ups of continuous fiber silicon carbide (SCS2) fiber/aluminum matrix composites were tested to assess fatigue mechanisms including stiffness loss when cycled below their respective fatigue limits. The lay-ups were 0 (sub 8), 0(sub 2)/ + or - 45 (sub 2s), 0/90 (sub 2s), and 0/ + or - 45/90 (subs). The data were compared with predictions from the author's previously published shakedown model which predicts fatigue-induced stiffness loss in metal matrix composites. A fifth lay-up, + or - 45 (sub 2s), was tested to compare shakedown and fatigue limits. The particular batch of silicon-carbide fibers tested in this program had a somewhat lower modulus (340 GPa) than expected and displayed poor bonding to the aluminum matrix. Good agreement was obtained between the stiffness loss model and the test data. The fatigue damage below the fatigue limit was primarily in the form of matrix cracking. The fatigue limit corresponded to the laminate shakedown for the + or - 45 (sub 2s) laminate.

Johnson, W. S.↗
NASA NTRS1986

Residual strength of five boron/aluminum laminates with crack-like notches after fatigue loading

Boron/aluminum specimens were made with crack-like slits in the center and with various proportions of 0 and + or - 45 deg plies. They were fatigue loaded and then fractured to determine their residual strengths. The fatigue loads were generally in the range of 60 to 80 percent of the static tensile strength of the specimen as determined from a previous study, and the stress ratio was .05. For virtually all of the specimens the fatigue loading was continued for 100,000 cycles. The specimens were radiographed after the fatigue loading to determine the nature of the fatigue damage. A few specimens were sectioned and examined in a scanning electron microscope after being radiographed in order to verify the interpretation of the radiographs and also to get a better insight into the nature of the fatigue damage. The results indicate that the fatiguing does not significantly affect the strength of the specimens tested. The results of the radiography and of the scanning electron microscopy indicate that the 45 deg plies suffer extensive damage in the form of split and broken fibers and matrix cracking in the vicinity of the ends of the split. By contrast, the only significant damage to the 0 deg plies was a single 0 deg matric crack growing from the ends of the slit and between the 0 deg fibers.

Simonds, R. A.↗
NASA NTRS1988

Extending Fatigue Lives Of Selected Alloys

Rejuvenating treatments demonstrated. Experiments with three alloys show some combinations of heat treatment, hot isostatic pressing, and/or surface treatment extends cycle lives of fatigue-damaged or fatigue-precracked specimens. For each alloy, tests were first done to determine "baseline" high-cycle-fatigue lives and high-cycle-fatigue failure modes over range of test conditions. Controlled fatigue damage introduced into specimens. Damaged specimens subjected to candidate rejuvenating treatments, then tested to failure. Effectiveness of each treatment judged by comparing fatigue data of damaged and treated specimens to "baseline" fatigue and data taken at same test conditions.

Matejczyk, D. E.↗
NASA NTRS1988

Fatigue testing and damage development in continuous fiber reinforced metal matrix composites

A general overview of the fatigue behavior of metal matrix composites (MMC) is presented. The first objective is to present experimental procedures and techniques for conducting a meaningful fatigue test to detect and quantify fatigue damage in MMC. These techniques include interpretation of stress-strain responses, acid etching of the matrix, edge replicas of the specimen under load, radiography, and micrographs of the failure surfaces. In addition, the paper will show how stiffness loss in continuous fiber reinforced metal matrix composites can be a useful parameter for detecting fatigue damage initiation and accumulation. Second, numerous examples of how fatigue damage can initiate and grow in various MMC are given. Depending on the relative fatigue behavior of the fiber and matrix, and the interface properties, the failure modes of MMC can be grouped into four categories: (1) matrix dominated, (2) fiber dominated, (3) self-similar damage growth, and (4) fiber/matrix interfacial failures. These four types of damage will be discussed and illustrated by examples with the emphasis on the fatigue of unnotched laminates.

Johnson, W. S.↗
NASA NTRS1988

The relationship between observed fatigue damage and life estimation models

Observations of the surface of laboratory specimens subjected to axial and torsional fatigue loadings has resulted in the identification of three damage fatigue phenomena: crack nucleation, shear crack growth, and tensile crack growth. Material, microstructure, state of stress/strain, and loading amplitude all influence which of the three types of fatigue damage occurs during a dominant fatigue life fraction. Fatigue damage maps are employed to summarize the experimental observations. Appropriate bulk stress/strain damage parameters are suggested to model fatigue damage for the dominant fatigue life fraction. Extension of the damage map concept to more complex loadings is presented.

Kurath, Peter↗
NASA NTRS1988

Creep-fatigue behavior of NiCoCrAlY coated PWA 1480 superalloy single crystals

Single crystal specimens of a Ni base superalloy, PWA 1480, with a low pressure plasma sprayed NiCoCrAlY coating were tested in various 0.1 Hz fatigue and creep fatigue cycles both at 1015 and 1050 C. Creep fatigue tests of the cp, pc, and cc types were conducted with various constant total strain ranges employing creep dwells at various constant stresses. Considerable cyclic softening occurred as was evidenced particularly by rapidly increasing creep rates in the creep fatigue tests. The cycle time in the creep fatigue tests typically decreased by more than 80 percent at 0.5 N sub f. Though cyclic life did correlate with delta epsilon sub in a better correlation existed with sub f for both the fatigue and creep fatigue tests, and poor correlations were observed with either sigma sub max or the average cycle time. A model containing both delta sigma and delta sigma (sub in), N sub f = alpha delta sigma (sub in) beta delta sigma gamma, with best fit values of sigma for each cycle type, but the same values of beta and gamam, was found to provide good correlations. Life lines were not greatly different among the cycle types, differing only by a factor of about three. The cp cycle life line was lowest for both test temperatures, however among the other three cycle types there was no consistent ranking. For all test types failure occurred predominately by multiple internal cracking originating at pores. The strong correlation of life with delta sigma may reflect a significant crack growth period in the life of the specimens.

Miner, R. V.↗
NASA NTRS1989

Fatigue crack growth study of SCS6/Ti-15-3 composite

A study was performed to determine the fatigue crack growth (FCG) behavior and the associated fatigue damage processes in a (0)(8) and (90)(8) oriented SCS(6)/Ti-15-3 composite. Companion testing (CT) was also done on identically processed Ti-15-3 unreinforced material. The active fatigue crack growth failure processes were very similar for both composite orientations tested. For both orientations, fatigue crack growth was along the fiber direction. It was found that the composite constituent most susceptible to fatigue damage was the interface region and in particular the carbon coating surrounding the fiber. The failure of the interface region lead to crack initiation and also strongly influenced the FCG behavior in this composite. The failure of the interface region was apparently driven by normal stresses perpendicular to the fiber direction. The FCG rates were considerably higher for the (90)(8) oriented CT specimens in comparison to the unreinforced material. This is consistent with the scenario in which the interface has lower fatigue resistance than the matrix, causing lower composite fatigue resistance. The FCG rates of the (0)(8) composite could not be directly compared to the (90)(8) composite but were shown to increase with an increase in the crack length.

Kantzos, Peter↗
NASA NTRS1990

A transient plasticity study and low cycle fatigue analysis of the Space Station Freedom photovoltaic solar array blanket

The Space Station Freedom photovoltaic solar array blanket assembly is comprised of several layers of materials having dissimilar elastic, thermal, and mechanical properties. The operating temperature of the solar array, which ranges from -75 to +60 C, along with the material incompatibility of the blanket assembly components combine to cause an elastic-plastic stress in the weld points of the assembly. The weld points are secondary structures in nature, merely serving as electrical junctions for gathering the current. The thermal mechanical loading of the blanket assembly operating in low earth orbit continually changes throughout each 90 min orbit, which raises the possibility of fatigue induced failure. A series of structural analyses were performed in an attempt to predict the fatigue life of the solar cell in the Space Station Freedom photovoltaic array blanket. A nonlinear elastic-plastic MSC/NASTRAN analysis followed by a fatigue calculation indicated a fatigue life of 92,000 to 160,000 cycles for the solar cell weld tabs. Additional analyses predict a permanent buckling phenomenon in the copper interconnect after the first loading cycle. This should reduce or eliminate the pulling of the copper interconnect on the joint where it is welded to the silicon solar cell. It is concluded that the actual fatigue life of the solar array blanket assembly should be significantly higher than the calculated 92,000 cycles, and thus the program requirement of 87,500 cycles (orbits) will be met. Another important conclusion that can be drawn from the overall analysis is that, the strain results obtained from the MSC/NASTRAN nonlinear module are accurate to use for low-cycle fatigue analysis, since both thermal cycle testing of solar cells and analysis have shown higher fatigue life than the minimum program requirement of 87,500 cycles.

Armand, Sasan C.↗
NASA NTRS1989

High temperature fatigue behavior of tungsten copper composites

The high temperature fatigue behavior of a 9 vol percent, tungsten fiber reinforced copper matrix composite was investigated. Load-controlled isothermal fatigue experiments at 260 and 560 C and thermomechanical fatigue (TMF) experiments, both in phase and out of phase between 260 and 560 C, were performed. The stress-strain response displayed considerable inelasticity under all conditions. Also, strain ratcheting was observed during all the fatigue experiments. For the isothermal fatigue and in-phase TMF tests, the ratcheting was always in a tensile direction, continuing until failure. The ratcheting during the out-of-phase TMF test shifted from a tensile direction to a compressive direction. This behavior was thought to be associated with the observed bulging and the extensive cracking of the out-of-phase specimen. For all cases, the fatigue lives were found to be controlled by damage to the copper matrix. Grain boundary cavitation was the dominant damage mechanism of the matrix. On a stress basis, TMF loading reduced lives substantially, relative to isothermal cycling. In-phase cycling resulted in the shortest lives, and isothermal fatigue at 260 C, the longest.

Verrilli, Michael J.↗
NASA NTRS1990

Open hole and post-impact compression fatigue of stitched and unstitched carbon/epoxy composites

The performance is studied of a stitched uniweave fabric composite and that of a toughened tape composite. The effects of stitching on compression fatigue life are addressed. Post impact compression fatigue and open hole fatigue tests were run on an AS4/3501-6 uniweave with stitching and a toughened IM7/8551-7 tape without stitching. Stitching was found to increase the thickness and consequently the weight of the composite material. The two materials were compared on an equal carbon content basis as well as on an equal weight basis. The excess thickness in the stitched uniweave composite was responsible for the lower fatigue life, on an equal carbon basis, compared to the toughened resin tape composite. Comparison of fatigue lives on an equal carbon content basis indicated that puncture or crimp type damage from stitching has very little effect on compression failure. Post impact fatigue test showed that although the damage area in the stitched uniweave composite was twice that of the toughened tape composite, the fatigue lives of the stitched composite were significantly longer than those of the toughened composite. Thus, it appears that the increase in thickness from stitching is much more of a penalty than crimped fibers or puncture type damage from stitching.

Portanova, M. A.↗