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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Split mandrel versus split sleeve coldworking: Dual methods for extending the fatigue life of metal structures

It is common practice to use split sleeve coldworking of fastener holes as a means of extending the fatigue life of metal structures. In search of lower manufacturing costs, the aerospace industry is examining the split mandrel (sleeveless) coldworking process as an alternative method of coldworking fastener holes in metal structures. The split mandrel process (SpM) significantly extends the fatigue life of metal structures through the introduction of a residual compressive stress in a manner that is very similar to the split sleeve system (SpSl). Since the split mandrel process is significantly less expensive than the split sleeve process and more adaptable to robotic automation, it will have a notable influence upon other new manufacture of metal structures which require coldworking a significant number of holes, provided the aerospace community recognizes that the resulting residual stress distributions and fatigue life improvement are the same for both processes. Considerable testing has validated the correctness of that conclusion. The findings presented in this paper represent the results of an extensive research and development program, comprising data collected from over 400 specimens fabricated from 2024-T3 and 7075-T651 aluminum alloys in varied configurations, which quantify the benefits (fatigue enhancement and cost savings) of automating a sleeveless coldworking system.

Rodman, Geoffrey A.↗
NASA NTRS1988

Improving the fatigue resistance of adhesive joints in laminated wood structures

The premature fatigue failure of a laminated wood/epoxy test beam containing a cross section finger joint was the subject of a multi-disciplinary investigation. The primary objectives were to identify the failure mechanisms which occurred during the finger joint test and to provide avenues for general improvements in the design and fabrication of adhesive joints in laminated wood structures.

Laufenberg, Theodore L.↗
NASA NTRS1993

Experimental research activities in dynamic response and sonic fatigue of hypersonic vehicle structures at NASA Langley Research Center

This paper presents an overview of experimental research activities being pursued at the NASA Langley Research Center for dynamic response and sonic fatigue of hypersonic vehicle structures. The capabilities of the principle test facility, the Thermal Acoustic Fatigue Apparatus, are first given. Results from recent dynamic response and sonic fatigue tests on candidate hypersonic vehicle structures are then presented.

Rizzi, Stephen A.↗
NASA NTRS2013

Modeling Delamination in Postbuckled Composite Structures Under Static and Fatigue Loads

The ability of the Abaqus progressive Virtual Crack Closure Technique (VCCT) to model delamination in composite structures was investigated for static, postbuckling, and fatigue loads. Preliminary evaluations were performed using simple Double Cantilever Beam (DCB) and Mixed-Mode Bending (MMB) specimens. The nodal release sequences that describe the propagation of the delamination front were investigated. The effect of using a sudden or a gradual nodal release was evaluated by considering meshes aligned with the crack front as well as misaligned meshes. Fatigue simulations were then performed using the Direct Cyclic Fatigue (DCF) algorithm. It was found that in specimens such as the DCB, which are characterized by a nearly linear response and a pure fracture mode, the algorithm correctly predicts the Paris Law rate of propagation. However, the Abaqus DCF algorithm does not consider different fatigue propagation laws in different fracture modes. Finally, skin/stiffener debonding was studied in an aircraft fuselage subcomponent in which debonding occurs deep into post-buckling deformation. VCCT was shown to be a robust tool for estimating the onset propagation. However, difficulties were found with the ability of the current implementation of the Abaqus progressive VCCT to predict delamination propagation within structures subjected to postbuckling deformations or fatigue loads.

Bisagni, Chiara↗
NASA NTRS1990

An evaluation of the pressure proof test concept for thin sheet 2024-T3

The concept of pressure proof testing of fuselage structures with fatigue cracks to insure structural integrity was evaluated from a fracture mechanics viewpoint. A generic analytical and experimental investigation was conducted on uniaxially loaded flat panels with crack configurations and stress levels typical of longitudinal lap splice joints in commercial transport aircraft fuselages. The results revealed that the remaining fatigue life after a proof test was longer than that without the proof test because of crack growth retardation due to increased crack closure. However, based on a crack length that is slightly less than the critical value at the maximum proof test stress, the minimum assured life or proof test interval must be no more than 550 pressure cycles for a 1.33 proof factor and 1530 pressure cycles for a 1.5 proof factor to prevent in-flight failures.

Dawicke, D. S.↗
NASA NTRS1991

An evaluation of the pressure proof test concept for 2024-T3 aluminium alloy sheet

The concept of pressure proof testing of fuselage structures with fatigue cracks to insure structural integrity was evaluated from a fracture mechanics viewpoint. A generic analytical and experimental investigation was conducted on uniaxially loaded flat panels with crack configurations and stress levels typical of longitudinal lap splice joints in commercial transport aircraft fuselages. The results revealed that the remaining fatigue life after a proof cycle was longer than that without the proof cycle because of crack growth retardation due to increased crack closure. However, based on a crack length that is slightly less than the critical value at the maximum proof stress, the minimum assured life or proof test interval must be no more than 550 pressure cycles for a 1.33 proof factor and 1530 pressure cycles for a 1.5 proof factor to prevent in-flight failures.

Dawicke, D. S.↗
NASA NTRS1990

An evaluation of the pressure proof test concept for thin sheet 2024-T3

The concept of pressure proof testing of fuselage structures with fatigue cracks to insure structural integrity was evaluated from a fracture mechanics viewpoint. A generic analytical and experimental investigation was conducted on uniaxially loaded flat panels with crack configurations and stress levels typical of longitudinal lap-splice joints in commercial transport aircraft fuselage. The results revealed that the remaining fatigue life after a proof test was longer than that without the proof test because of crack growth retardation due to increased crack closure. However, based on a crack length that is slightly less than the critical value at the maximum proof test stress, the minimum assured life or proof test interval must be no more than 550 pressure cycles for a 1.33 proof factor and 1530 pressure cycles for a 1.5 proof factor to prevent in-flight failures.

Dawicke, D. S.↗
NASA NTRS2010

The Effect of a Non-Gaussian Random Loading on High-Cycle Fatigue of a Thermally Post-Buckled Structure

High-cycle fatigue of an elastic-plastic beam structure under the combined action of thermal and high-intensity non-Gaussian acoustic loadings is considered. Such loadings can be highly damaging when snap-through motion occurs between thermally post-buckled equilibria. The simulated non-Gaussian loadings investigated have a range of skewness and kurtosis typical of turbulent boundary layer pressure fluctuations in the vicinity of forward facing steps. Further, the duration and steadiness of high excursion peaks is comparable to that found in such turbulent boundary layer data. Response and fatigue life estimates are found to be insensitive to the loading distribution, with the minor exception of cases involving plastic deformation. In contrast, the fatigue life estimate was found to be highly affected by a different type of non-Gaussian loading having bursts of high excursion peaks.

Rizzi, Stephen A.↗
NASA NTRS1990

A model for predicting damage induced fatigue life of laminated composite structural components

This paper presents a model for predicting the life of laminated composite structural components subjected to fatigue induced microstructural damage. The model uses the concept of continuum damage mechanics, wherein the effects of microcracks are incorporated into a damage dependent lamination theory instead of treating each crack as an internal boundary. Internal variables are formulated to account for the effects of both matrix cracks and internal delaminations. Evolution laws for determining the damage variables as functions of ply stresses are proposed, and comparisons of predicted damage evolution are made to experiment. In addition, predicted stiffness losses, as well as ply stresses are shown as functions of damage state for a variety of stacking sequences.

Allen, David H.↗
NASA NTRS1974

Strain-cycling fatigue behavior of ten structural metals tested in liquid helium (4 K), in liquid nitrogen (78 K), and in ambient air (300 K)

Strain-cycling fatigue behavior of 10 different structural alloys and metals was investigated in liquid helium (4 K), in liquid nitrogen (78 K), and in ambient air (300 K). At high cyclic lives, fatigue resistance increased with decreasing temperature for all the materials investigated. At low cyclic lives, fatigue resistance generally decreased with decreasing temperature for the materials investigated. Only for Inconel 718 did fatigue resistance increase with decreasing temperature over the entire life range investigated. Comparison of the experimental fatigue behavior with that predicted by the Manson method of universal slopes showed that the fatigue behavior of these materials can be predicted for cryogenic temperatures by using material tensile properties obtained at those same temperatures.

Nachtigall, A. J.↗
NASA NTRS1976

Optimal periodic proof test based on cost-effective and reliability criteria

An exploratory study for the optimization of periodic proof tests for fatigue-critical structures is presented. The optimal proof load level and the optimal number of periodic proof tests are determined by minimizing the total expected (statistical average) cost, while the constraint on the allowable level of structural reliability is satisfied. The total expected cost consists of the expected cost of proof tests, the expected cost of structures destroyed by proof tests, and the expected cost of structural failure in service. It is demonstrated by numerical examples that significant cost saving and reliability improvement for fatigue-critical structures can be achieved by the application of the optimal periodic proof test. The present study is relevant to the establishment of optimal maintenance procedures for fatigue-critical structures.

Yang, J.-N.↗
NASA NTRSDate not supplied

Progressive Damage Mechanisms in Thick 3D Composite Structures under Static and Fatigue Loading

The objective of this work is to demonstrate validation of predictive capabilities capturing both damage initiation and growth under static and fatigue loading in geometrically complex three-dimensional (3D) composite structures. The composite structure of interest in the present study resembles a 3D blade spar used in rotorcraft applications. Under both static and fatigue testing, it was shown that damage initiated at the same ply-drop location(s) and grew in the axial as well as hoop directions. CompDam – a state-of-the-art Progressive Damage Analysis (PDA) tool was used to achieve successful correlations of metrics including load-displacement responses and damage modes.

progressive damage analysis↗
NASA NTRS1972

Establishment of design criteria for acceptable failure modes and fail safe considerations for the space shuttle structural system

Investigations of fatigue life, and safe-life and fail-safe design concepts as applied to space shuttle structure are summarized. The results are evaluated to select recommended structural design criteria to provide assurance that premature failure due to propagation of undetected crack-like defects will not occur during shuttle operational service. The space shuttle booster, GDC configuration B-9U, is selected as the reference vehicle. Structural elements used as basis of detail analyses include wing spar caps, vertical stabilizer skins, crew compartment skin, orbiter support frame, and propellant tank shell structure. Fatigue life analyses of structural elements are performed to define potential problem areas and establish upper limits of operating stresses. Flaw growth analyses are summarized in parametric form over a range of initial flaw types and sizes, operating stresses and service life requirements. Service life of 100 to 500 missions is considered.

Westrup, R. W.↗
NASA NTRS2019

Damage Mechanisms in Tapered Composite Structures Under Static and Fatigue Loading

In this work an integrated computational/experimental approach was developed to validate the predictive capabilities of State-of-the-Art (SoA) Progressive Damage Analysis (PDA) methods and tools. Specifically, a tapered composite structure incorporating ply-drops typical in the aerospace industry to spatially vary structural thickness was tested under static tension and cyclic tension fatigue loads. The data acquired from these tests included quantitative metrics such as pre-peak stiffness, peak load, location of delamination damage onset, and growth of delaminations as functions of applied static and fatigue loads. It was shown that the PDA tools were able to predict the pre-peak stiffness and peak load within ±10% of experimental average, thereby meeting and exceeding the pre-defined success criteria. Additionally, it was shown that the PDA tools were able to accurately predict the location of delamination onset and satisfactorily predict delamination growth under static tension loading. Overall, good correlations were achieved between modeling and experiments.

Rao, Prabhakar M.↗
NASA NTRS2007

On the Use of 3dB Qualification Margin for Structural Parts on Expendable Launch Vehicles

The standard random vibration qualification test used for Expendable Launch Vehicle components is Maximum Predicted Environment (MPE) + 6dB for a duration of 4 times the service life of the part. This can be a severe qualification test for these fatigue-sensitive structures. This paper uses flight data from several launch vehicles to establish that reducing the qualification approach to MPE+3dB for the duration of the peak environment (1x life) is valid for fatigue-sensitive structural components. Items that can be classified as fatigue-sensitive are probes, ducts, tubing, bellows, hoses, and any non-functional structure. Non-functional structure may be flight critical or carry fluid, but it cannot include any moving parts or electronics. This reduced qualification approach does not include primary or secondary structure which would be exclusively designed by peak loads, either transient or quasi-static, that are so large and of so few cycles as to make fatigue a moot point.

Yunis, Isam↗
NASA NTRS1995

Empirical modeling of environment-enhanced fatigue crack propagation in structural alloys for component life prediction

This research aims to develop the methods and understanding needed to incorporate time and loading variable dependent environmental effects on fatigue crack propagation (FCP) into computerized fatigue life prediction codes such as NASA FLAGRO (NASGRO). In particular, the effect of loading frequency on FCP rates in alpha + beta titanium alloys exposed to an aqueous chloride solution is investigated. The approach couples empirical modeling of environmental FCP with corrosion fatigue experiments. Three different computer models have been developed and incorporated in the DOS executable program. UVAFAS. A multiple power law model is available, and can fit a set of fatigue data to a multiple power law equation. A model has also been developed which implements the Wei and Landes linear superposition model, as well as an interpolative model which can be utilized to interpolate trends in fatigue behavior based on changes in loading characteristics (stress ratio, frequency, and hold times).

Richey, Edward, III↗
NASA NTRS1989

Application of fracture mechanics and half-cycle theory to the prediction of fatigue life of aerospace structural components

The service life of aircraft structural components undergoing random stress cycling was analyzed by the application of fracture mechanics. The initial crack sizes at the critical stress points for the fatigue crack growth analysis were established through proof load tests. The fatigue crack growth rates for random stress cycles were calculated using the half-cycle method. A new equation was developed for calculating the number of remaining flights for the structural components. The number of remaining flights predicted by the new equation is much lower than that predicted by the conventional equation. This report describes the application of fracture mechanics and the half-cycle method to calculate the number of remaining flights for aircraft structural components.

Ko, William L.↗