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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 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Excitation, response, and fatigue life estimation methods for the structural design of externally blown flaps

Means for predicting the fluctuating pressures acting on externally blown flap surfaces are developed on the basis of generalizations derived from non-dimensionalized empirical data. Approaches for estimation of the fatigue lives of skin-stringer and honeycomb-core sandwich flap structures are derived from vibration response analyses and panel fatigue data. Approximate expressions for fluctuating pressures, structural response, and fatigue life are combined to reveal the important parametric dependences. The two-dimensional equations of motion of multi-element flap systems are derived in general form, so that they can be specialized readily for any particular system. An introduction is presented of an approach to characterizing the excitation pressures and structural responses which makes use of space-time spectral concepts and promises to provide useful insights, as well as experimental and analytical savings.

Ungar, E. E.↗
NASA NTRS2007

Fatigue Life of Haynes 188 Superalloy in Direct Connect Combustor Durability Rig

The Direct Connect Combustor Durability Rig (DCR) will provide NASA a flexible and efficient test bed to demonstrate the durability of actively cooled scramjet engine structure, static and dynamic sealing technologies, and thermal management techniques. The DCR shall be hydrogen fueled and cooled, and test hydrogen coolded structural panels at Mach 5 and 7. Actively cooled Haynes 188 superalloy DCR structural panels exposed to the combustion environment shall have electrodischarge machined (EDM) internal cooling holes with flowing liquid hydrogen. Hydrogen combustion could therefore produce severe thermal conditions that could challenge low cycle fatigue durability of this material. The objective of this study was to assess low cycle fatigue capability of Haynes 188 for DCR application. Tests were performed at 25 and 650 C, in hydrogen and helium environments, using specimens with low stress ground (LSG) and electro-discharge machined (EDM) surface finish. Initial fatigue tests in helium and hydrogen indicate the low cycle fatigue life capability of Haynes 188 in hydrogen appears quite satisfactory for the DCR application. Fatigue capability did not decrease with increasing test temperature. Fatigue capability also did not decrease with EDM surface finish. Failure evaluations indicate retention of ductility in all conditions. Additional tests are planned to reconfirm these positive trends.

Gabb, TIm↗
NASA NTRS2015

Effect of Buckling Modes on the Fatigue Life and Damage Tolerance of Stiffened Structures

The postbuckling response and the collapse of composite specimens with a co-cured hat stringer are investigated experimentally and numerically. These specimens are designed to evaluate the postbuckling response and the effect of an embedded defect on the collapse load and the mode of failure. Tests performed using controlled conditions and detailed instrumentation demonstrate that the damage tolerance, fatigue life, and collapse loads are closely tied with the mode of the postbuckling deformation, which can be different between two nominally identical specimens. Modes that tend to open skin/stringer defects are the most damaging to the structure. However, skin/stringer bond defects can also propagate under shearing modes. In the proposed paper, the effects of initial shape imperfections on the postbuckling modes and the interaction between different postbuckling deformations and the propagation of skin/stringer bond defects under quasi-static or fatigue loads will be examined.

Davila, Carlos G.↗
NASA NTRS2009

Development of a Composite Delamination Fatigue Life Prediction Methodology

Delamination is one of the most significant and unique failure modes in composite structures. Because of a lack of understanding of the consequences of delamination and the inability to predict delamination onset and growth, many composite parts are unnecessarily rejected upon inspection, both immediately after manufacture and while in service. NASA Langley is leading the efforts in the U.S. to develop a fatigue life prediction methodology for composite delamination using fracture mechanics. Research being performed to this end will be reviewed. Emphasis will be placed on the development of test standards for delamination characterization, incorporation of approaches for modeling delamination in commercial finite element codes, and efforts to mature the technology for use in design handbooks and certification documents.

OBrien, Thomas K.↗
NASA NTRS1984

Oxidation and low cycle fatigue life prediction

When a metallic material is exposed to a high temperature in an ambient atmosphere, oxidation takes place on the metallic surface. The formed oxides (both surface and grain boundary oxides) are mechanically brittle so that if the stress is high enough the oxides will be cracked. The grain boundary oxide formation in TAZ-8A nickel-base superalloy was studied. The effect of oxide crack nucleus on low cycle fatigue life will be analyzed. The TAZ-8A was subjected to high temperature oxidation tests in air under the stress-free condition. The oxidation temperatures were 600, 800, and 1000 C. The oxidation time varies from 10 to 1000 hours.

Oshida, Y.↗
NASA NTRS2015

Microstructural Influence on Deformation and Fatigue Life of Composites Using the Generalized Method of Cells

A fully coupled deformation and damage approach to modeling the response of composite materials and composite laminates is presented. It is based on the semi-­‐analytical generalized method of cells (GMC) micromechanics model as well as its higher fidelity counterpart, HFGMC, both of which provide closed-form constitutive equations for composite materials as well as the micro scale stress and strain fields in the composite phases. The provided constitutive equations allow GMC and HFGMC to function within a higher scale structural analysis (e.g., finite element analysis or lamination theory) to represent a composite material point, while the availability of the micro fields allow the incorporation of lower scale sub­‐models to represent local phenomena in the fiber and matrix. Further, GMC's formulation performs averaging when applying certain governing equations such that some degree of microscale field accuracy is surrendered in favor of extreme computational efficiency, rendering the method quite attractive as the centerpiece in a integrated computational material engineering (ICME) structural analysis; whereas HFGMC retains this microscale field accuracy, but at the price of significantly slower computational speed. Herein, the sensitivity of deformation and the fatigue life of graphite/epoxy PMC composites, with both ordered and disordered microstructures, has been investigated using this coupled deformation and damage micromechanics based approach. The local effects of fiber breakage and fatigue damage are included as sub‐models that operate on the microscale for the individual composite phases. For analysis of laminates, classical lamination theory is employed as the global or structural scale model, while GMC/HFGMC is embedded to operate on the microscale to simulate the behavior of the composite material within each laminate layer. A key outcome of this study is the statistical influence of microstructure and micromechanics idealization (GMC or HFGMC) on the overall accuracy of unidirectional and laminated composite deformation and fatigue response.

Microscale Models↗
NASA NTRS1972

Effect of residual stresses induced by prestressing on rolling element fatigue life

A mechanical prestress cycle suitable to induce compressive stress beneath the surface of the inner race of radially loaded 207-size bearings was determined. Compressive residual stress in excess 0.69 x 10 to the 9th power N/sq m (100,000 psi), as measured by X-ray diffraction, were induced at the depth of maximum shearing stress. The prestress cycle consisted of running the bearings for 25 hours at 2750 rpm at a radial load which produced a maximum Hertz stress of 3.3 x 10 to the 9th power N/sq m (480,000 psi) at the contact of the inner race and the heaviest loaded ball. Bearings subjected to this prestress cycle and subsequently fatigue tested gave a 10 percent fatigue life greater than twice that of a group of baseline bearings.

Parker, R. J.↗
NASA NTRS1988

Surface fatigue life of CBN and vitreous ground carburized and hardened AISI 9310 spur gears

Spur gear surface endurance tests were conducted to investigate CBN ground AISI 9310 spur gears for use in aircraft applications, to determine their endurance characteristics and to compare the results with the endurance of standard vitreous ground AISI 9310 spur gears. Tests were conducted with VIM-VAR AISI 9310 carburized and hardened gears that were finish ground with either CBN or vitreous grinding methods. Test conditions were an inlet oil temeprature of 320 K (116 F), an outlet oil temperature of 350 K (170 F), a maximum Hertz stress of 1.71 GPa (248 ksi), and a speed of 10,000 rpm. The CBN ground gears exhibited a surface fatigue life that was slightly better than the vitreous ground gears. The subsurface residual stress of the CBN ground gears was approximately the same as that for the standard vitreous ground gears for the CBN grinding method used.

Townsend, Dennis P.↗
NASA NTRS2006

Fatigue Life Methodology for Tapered Hybrid Composite Flexbeams

Nonlinear-tapered flexbeam specimens from a full-size composite helicopter rotor hub flexbeam were tested under combined constant axial tension and cyclic bending loads. Two different graphite/glass hybrid configurations tested under cyclic loading failed by delamination in the tapered region. A 2-D finite element model was developed which closely approximated the flexbeam geometry, boundary conditions, and loading. The analysis results from two geometrically nonlinear finite element codes, ANSYS and ABAQUS, are presented and compared. Strain energy release rates (G) associated with simulated delamination growth in the flexbeams are presented from both codes. These results compare well with each other and suggest that the initial delamination growth from the tip of the ply-drop toward the thick region of the flexbeam is strongly mode II. The peak calculated G values were used with material characterization data to calculate fatigue life curves for comparison with test data. A curve relating maximum surface strain to number of loading cycles at delamination onset compared well with the test results.

urri, Gretchen B.↗
NASA NTRS2002

Fatigue Life Analysis of Tapered Hybrid Composite Flexbeams

Nonlinear-tapered flexbeam laminates from a full-size composite helicopter rotor hub flexbeam were tested under combined constant axial tension and cyclic bending loads. The two different graphite/glass hybrid configurations tested under cyclic loading failed by delamination in the tapered region. A 2-D finite element model was developed which closely approximated the flexbeam geometry, boundary conditions, and loading. The analysis results from two geometrically nonlinear finite element codes, ANSYS and ABAQUS, are presented and compared. Strain energy release rates (G) obtained from the above codes using the virtual crack closure technique (VCCT) at a resin crack location in the flexbeams are presented for both hybrid material types. These results compare well with each other and suggest that the initial delamination growth from the resin crack toward the thick region of the flexbeam is strongly mode II. The peak calculated G values were used with material characterization data to calculate fatigue life curves and compared with test data. A curve relating maximum surface strain to number of loading cycles at delamination onset compared reasonably well with the test results.

Murri, Gretchen B.↗
NASA NTRS1990

Airframe structural optimization for maximum fatigue life

A methodology is outlined for optimization of airframe structures under dynamic constraints to maximize service life of specified fatigue-critical components. For practical airframe structures, this methodology describes the development of sensitivity analysis and computational procedures for constraints on the steady-state dynamic response displacements and stresses. Strain energy consideration is used for selection of structural members for modification. Development of a design model and its relation to an analysis model, as well as ways to reduce the dimensionality of the problem via approximation concepts, are described. This methodology is demonstrated using an elastic stick model for the MH-53J helicopter to show service life improvements of the hinge fold region.

D. P. Schrage↗
NASA NTRS1942

Improvement of fatigue life of an aluminum alloy by overstressing

Fatigue tests were made on some 1.375-inch-diameter and 0.300-inch diameter specimens of a 17S-T aluminum alloy rod. One test of a large specimen was run continuously to failure at a maximum stress of 22,000 pounds per square inch. In two other tests of large specimens, thin surface layers were removed periodically until failure occurred. The same nominal maximum stress of 22,000 pounds per square inch was used throughout the two tests and the load on the fatigue machine was lowered accordingly after the removal of each surface layer. As each test progressed the stress in the metal of the final surface area therefore was increased after the removal of each surface layer. Because of the stresses used, this metal was overstressed, that is, stressed above its endurance limit. All the remaining specimens were subjected to similar over- stressing conditions but no metal was removed and a low initial stress was increased periodically to a final maximum value of 22,000 pounds per square inch as each test progressed. It was found that the fatigue resistance of 17S-T aluminum alloy can be increased by moderate overstressing. Apparently the increase in fatigue lift obtained in the tests of specimens from which layers were removed was the result of overstressing rather than from the removal of damaged surface layers.

STRICKLEY G W↗
NASA NTRS1993

Fatigue life and crack growth prediction methodology

The capabilities of a plasticity-induced crack-closure model and life-prediction code to predict fatigue crack growth and fatigue lives of metallic materials are reviewed. Crack-tip constraint factors, to account for three-dimensional effects, were selected to correlate large-crack growth rate data as a function of the effective-stress-intensity factor range (delta(K(sub eff))) under constant-amplitude loading. Some modifications to the delta(K(sub eff))-rate relations were needed in the near threshold regime to fit small-crack growth rate behavior and endurance limits. The model was then used to calculate small- and large-crack growth rates, and in some cases total fatigue lives, for several aluminum and titanium alloys under constant-amplitude, variable-amplitude, and spectrum loading. Fatigue lives were calculated using the crack growth relations and microstructural features like those that initiated cracks. Results from the tests and analyses agreed well.

Newman, J. C., Jr.↗
NASA NTRS1994

Fatigue life and crack growth prediction methodology

This paper reviews the capabilities of a plasticity-induced crack-closure model and life-prediction code to predict fatigue crack growth and fatigue lives of metallic materials. Crack-tip constraint factors, to account for three-dimensional effects, were selected to correlate large-crack growth rate data as a function of the effective stress-intensity factor range (Delta K(sub eff)) under constant amplitude loading. Some modifications to the Delta K(sub eff)-rate relations were needed in the near threshold regime to fit small-crackgrowth rate behavior and endurance limits. The model was then used to calculate small- and large-crack growth rates, and in some cases total fatigue lives, for several aluminum and titanium alloys under constant-amplitude, variable-amplitude, and spectrum loading. Fatigue lives were calculated using the crack-growth relations and microstructural features like those that initiated cracks. Results from the tests and analyses agreed well.

Newman, J. C., Jr.↗
NASA NTRS2003

FLAPS (Fatigue Life Analysis Programs): Computer Programs to Predict Cyclic Life Using the Total Strain Version of Strainrange Partitioning and Other Life Prediction Methods

This manual presents computer programs FLAPS for characterizing and predicting fatigue and creep-fatigue resistance of metallic materials in the high-temperature, long-life regime for isothermal and nonisothermal fatigue. The programs use the Total Strain version of Strainrange Partitioning (TS-SRP), and several other life prediction methods described in this manual. The user should be thoroughly familiar with the TS-SRP and these life prediction methods before attempting to use any of these programs. Improper understanding can lead to incorrect use of the method and erroneous life predictions. An extensive database has also been developed in a parallel effort. The database is probably the largest source of high-temperature, creep-fatigue test data available in the public domain and can be used with other life-prediction methods as well. This users' manual, software, and database are all in the public domain and can be obtained by contacting the author. The Compact Disk (CD) accompanying this manual contains an executable file for the FLAPS program, two datasets required for the example problems in the manual, and the creep-fatigue data in a format compatible with these programs.

Arya, Vinod K.↗
NASA NTRS1968

Predicting fatigue life of metal bellows

Classical method of presenting fatigue data in plots of alternating stress vs number of deflection cycles is applied to bellows formed of various metals, including corrosion-resistant steel, nickel alloys, and aluminum alloys. The expected life of a new bellows design can then be determined before fabrication and testing.

Daniels, C. M.↗
NASA NTRS1985

Creep fatigue life prediction for engine hot section materials (isotropic)

The first two years of a two-phase program aimed at improving the high temperature crack initiation life prediction technology for gas turbine hot section components are discussed. In Phase 1 (baseline) effort, low cycle fatigue (LCF) models, using a data base generated for a cast nickel base gas turbine hot section alloy (B1900+Hf), were evaluated for their ability to predict the crack initiation life for relevant creep-fatigue loading conditions and to define data required for determination of model constants. The variables included strain range and rate, mean strain, strain hold times and temperature. None of the models predicted all of the life trends within reasonable data requirements. A Cycle Damage Accumulation (CDA) was therefore developed which follows an exhaustion of material ductility approach. Material ductility is estimated based on observed similarities of deformation structure between fatigue, tensile and creep tests. The cycle damage function is based on total strain range, maximum stress and stress amplitude and includes both time independent and time dependent components. The CDA model accurately predicts all of the trends in creep-fatigue life with loading conditions. In addition, all of the CDA model constants are determinable from rapid cycle, fully reversed fatigue tests and monotonic tensile and/or creep data.

Moreno, Vito↗