Spectrum Loading Conditions Research Articles

Cycle-by-cycle crack closure measurements and fatigue crack growth in CT specimen made of 7075-T7351 aluminium

In this study, an advanced piezoelectric strain sensor is applied to measure the crack opening loads of a compact tension specimen with a fatigue crack under variable amplitude block loading. The crack opening loads are subsequently utilised to modify the stress intensity factor range by taking into account the proportion of each load cycle for which the fatigue crack remained closed. This effective stress intensity factor is applied to collapse the growth rate curve, which initially displays a strong dependence on stress ratio, into a single master curve where the only variable governing crack extension is the effective stress intensity factor. The findings of this study provide strong evidence for the crack closure concept, which is important for the development of algorithms which can compress/truncate large spectrums. These algorithms rely on a detailed understanding of how much the crack is open or closed during each load cycle. Furthermore, the results indicate that the piezoelectric sensor could be used to study crack closure under realistic spectrum loading conditions, which will be the subject of future investigations.

Fatigue crack growth of a railway wheel steel and fatigue life prediction under spectrum loading conditions

In this paper, fatigue life and crack growth retardation of railway wheel specimens due to spectrum loading have been investigated. For this purpose, numerical and experimental methods are used to investigate the behavior of crack growth in the wheel specimens. The crack propagation is examined by using the stress intensity factor (SIF) values achieved from boundary element (BE) analyses. The non‐linear Kujawski‐Ellyin (KE) model has been adopted for the crack propagation under small‐scale yielding (SSY) conditions. The periodic overloads have remained constant and the effect of overload ratio (OLR) and occurrence ratios (OCR) on the fatigue life has been investigated.

A quantitative assessment of the impact of corrosion on fatigue life of aircraft components

Management of corrosion, and repair and restoration of corrosion‐affected structures pose a daunting challenge to the aircraft industry. This paper presents analytical studies to examine and quantify the effects of corrosion of Al-alloys on the fatigue life of aircraft structures. Two scenarios are considered: (i) fatigue of a pre-corroded material and (ii) fatigue in a corrosive environment. The corrosion damage is idealized using standard crack models, and cycle-by-cycle fatigue crack growth analysis is carried out for an aircraft horizontal stabilizer under spectrum loading conditions. A few aircraft operational scenarios are considered, and fatigue life calculations presented. The zone which experiences 18% lower stress than the maximum stress zone is shown to turn fatigue critical if vulnerable to corrosion. The results point to the risk posed by corroded zones if missed while focusing attention on high stress zones. The computations indicate that an exposure of even 25% of operational life to corrosive environment can reduce life by 40–55%. The study quantifies the significance of corrosion while arriving at design safe life, remaining useful life and maintenance intervals.

Fatigue life prediction based on equivalent initial flaw size for Al-Li alloy 2297 under spectrum loading

Fatigue life prediction methods based on the Equivalent Initial Flaw Size (EIFS) concept under variable amplitude (VA) spectrum loading conditions were studied in this paper. Fatigue tests of Al-Li alloy 2297 under constant amplitude (CA) and VA loading cases were carried out. Based on the comparison of predictions and test results, different ways to determine the EIFSs and whether the load-interaction effects should be considered in VA fatigue life prediction were discussed to recommend their applicable cases. A good agreement is observed between EIFS model predictions and experimental data, especially in case of high maximum stresses in VA loading.

Random spectrum loading of dental implants: An alternative approach to functional performance assessment

The fatigue performance of dental implants is usually assessed on the basis of cyclic S/N curves. This neither provides information on the anticipated service performance of the implant, nor does it allow for detailed comparisons between implants unless a thorough statistical analysis is performed, of the kind not currently required by certification standards.The notion of endurance limit is deemed to be of limited applicability, given unavoidable stress concentrations and random load excursions, that all characterize dental implants and their service conditions.We propose a completely different approach, based on random spectrum loading, as long used in aeronautical design.The implant is randomly loaded by a sequence of loads encompassing all load levels it would endure during its service life. This approach provides a quantitative and comparable estimate of its performance in terms of lifetime, based on the very fact that the implant will fracture sooner or later, instead of defining a fatigue endurance limit of limited practical application.Five commercial monolithic Ti-6Al-4V implants were tested under cyclic, and another 5 under spectrum loading conditions, at room temperature and dry air. The failure modes and fracture planes were identical for all implants. The approach is discussed, including its potential applications, for systematic, straightforward and reliable comparisons of various implant designs and environments, without the need for cumbersome statistical analyses.It is believed that spectrum loading can be considered for the generation of new standardization procedures and design applications.

Mechanism of Friction and Wear in MoS2 and ZDDP/F-PTFE Greases under Spectrum Loading Conditions

Two different greases formulated using MoS2 and a combination of ZDDP and functionalized PTFE (F-PTFE) were examined under spectrum loading conditions where loads, frequency, and duration of the steps were treated as variables. Combination of ZDDP and F-PTFE were synergistic resulting in a significant reduction in the wear and friction under spectrum loading condition. Decreasing the time step during the ramp up and ramp down cycles resulted in larger wear for the grease containing MoS2 particles in comparison to ZDDP/F-PTFE in grease. The tribofilm formed on the surface was analyzed using various characterization techniques like SEM, EDS, and Stereo Optical Microscopy. Tribofilms from MoS2 additives had extensive amounts of abrasive and adhesive wear and showed the formation of MoS2 on the surface on the other hand the tribofilms from ZDDP/F-PTFE had smaller amounts of severe wear and exhibited patchy tribofilms of Zn-phosphates as well as sulfides of Zn and Fe.

Role of MoS2morphology on wear and friction under spectrum loading conditions

AbstractOne of the ways by which grease is evaluated is by using a four‐ball wear test using ASTM D2266. However, actual applications may require bearings to be subjected to spectrum loading conditions. This study focuses on using ball milling to mitigate the wear from sharp edges in the MoS2particles. Two different blends of greases were formulated using MoS2in the as‐received state (unmilled) and milled MoS2; they were tested under spectrum loading conditions where the load and frequency of the tests were treated as variables. It was found that ball milling of the MoS2significantly reduces the wear under spectrum loading condition both for ramp‐up and ramp‐down conditions. It was also shown that shortening the time step for both the ramp‐up and ramp‐down cycles resulted in larger wear for unmilled MoS2particles in comparison with milled MoS2particles in grease. The milling process did not play a significant role when frequency of the test was either ramped up or down. Copyright © 2015 John Wiley & Sons, Ltd.

Load and temperature interaction modeling of fatigue crack growth in a Ni-base superalloy

An approach was developed to predict the thermo-mechanical fatigue crack growth rates under typical gas turbine engine spectrum loading conditions. The material studied in the development of this model was a polycrystalline superalloy, Inconel 100. Load interaction effects were determined to have a major effect on the crack growth life. A yield zone load interaction life prediction model was modified to include temperature dependent properties. Multiple overload effects were included in the model to incorporate enhanced retardation compared to single overload retardation behavior. Temperature interaction effects were included and proved to be very important because of the wide temperature ranges to which turbine engine components are subjected. The effects of oxidation and temperature changes were accounted for in the model by accelerating crack growth in regions that had been previously affected by elevated temperatures. Experimental data of isolated, first order effects were used to calibrate and verify the model. Temperature dependent mechanical properties were determined and were essential in the model’s development. Parametric studies were performed using this model to assess the sensitivity of specific crack growth variables on life predictions.

Estimating crack growth in rotorcraft structures subjected to mission load spectrum

This paper is an extension of the work originally reported on the international conference of experimental mechanics. In this extended work, numerical and experimental results are presented from a project aimed at predicting the fatigue life of a rotorcraft structural component subjected to flight load spectrum. The structural component is a riveted joint used in cabin frame cap splices of several civilian and military helicopters modeled herein as a lap-joined nested angle assembly. This component is fatigue sensitive due to the highly cyclic and vibratory nature of a rotorcraft mission load spectrum and as such accurate prediction of its fatigue life is of major importance in the design cycle. In this work, numerical studies are conducted using constant amplitude and mission spectrum loads. The rate of fatigue crack growth for through-the-thickness crack initiated from the critical fastener hole is computed using 2D standard and weight function models with the crack plane stress field obtained from 3D finite element analysis. Effect of load interaction due to tensile overload is included using strip-yield retardation model and the effect of surface shear arising from contact friction between the plates and the fasteners is studied using parametric models. Finally, results of the numerical simulations are compared with representative experimental data obtained under similar spectrum loading condition.

Improvement of damage tolerance of laser beam welded stiffened panels for airframes via local engineering

Damage tolerance of an aerospace grade aluminum alloy was studied using a new design philosophy in skin and stringer geometries. Systematic thickness variations (crenellations) were introduced onto the skin and stringers of the laser beam welded (LBW) stiffened Al2139-T8 large center cracked flat panels to modify the stress intensity factor (SIF) distribution and hence to improve fatigue life. Fatigue crack propagation (FCP) tests (on panels with crenellations) with crack growing perpendicular to the welded stringers were conducted under constant amplitude and spectrum loading conditions. Results were compared with the “classical” LBW stiffened panels (with no crenellations) having equal weight and tested under the same conditions. The new panel design with crenellations showed substantially longer fatigue lives under constant amplitude loading. This gain significantly improved under spectrum (MINI-TWIST) loading fatigue tests. This paper presents the first FCP test results of a comprehensive ongoing program which investigates the efficiency of component design with crenellations to improve damage tolerance behavior of welded Al-alloy and steel structures. Issues including microstructural examinations, numerical investigations, fitness-for-service (FFS) analysis and residual strength aspects of this program will be topics of another communication.

Effect of loading parameters on the fatigue behavior of impact damaged composite laminates

The long-term mechanical fatigue of AS4/3501-6 graphite/epoxy quasi-isotropic laminates has been investigated to determine the influence of loading parameters on impact-induced delamination growth during constant amplitude, block, and spectrum fatigue loading. From the observed fatigue damage growth it is suggested that conservative damage tolerance criteria in the presence of barely visible impact damage should be based on static loading tests and a limited number of fatigue tests to determine the load levels that initiate delamination growth. For the type of impact damage and laminate configuration investigated in this study the fatigue limit is found to be 60% CSAI (compression strength after impact) in constant-amplitude compression-compression loading, and 65% CSAI under the spectrum loading conditions. Furthermore, the results of spectrum loading tests suggest that the durability testing can be accelerated by omitting from the standardized test sequence load levels in which the stress range is less than 20% CSAI.

SMALL CRACK GROWTH AND FATIGUE LIFE PREDICTIONS FOR HIGH‐STRENGTH ALUMINIUM ALLOYS: PART I—EXPERIMENTAL AND FRACTURE MECHANICS ANALYSIS

The small crack effect was investigated in two high‐strength aluminium alloys: 7075‐T6 bare and LC9cs clad alloy. Both experimental and analytical investigations were conducted to study crack initiation and growth of small cracks. In the experimental program, fatigue tests, small crack and large crack tests were conducted under constant amplitude and Mini‐TWIST spectrum loading conditions. A pronounced small crack effect was observed in both materials, especially for the negative stress ratios. For all loading conditions, most of the fatigue life of the SENT specimens was shown to be crack propagation from initial material defects or from the cladding layer. In the analysis program, three‐dimensional finite element and weight function methods were used to determine stress intensity factors and to develop SIF equations for surface and corner cracks at the notch in the SENT specimens. A plasticity‐induced crack‐closure model was used to correlate small and large crack data, and to make fatigue life predictions. Predicted crack‐growth rates and fatigue lives agreed well with experiments. A total fatigue life prediction method for the aluminium alloys was developed and demonstrated using the crack‐closure model.

A study of spectrum fatigue crack propagation in two aluminum alloys—I. Spectrum simplification

A study of the fatigue crack propagation behavior of two commercial Al alloys was undertaken using spectrum loading conditions characteristic of those encountered at critical locations in high performance fighter aircraft. Tension-dominated (TD) and tension-compression (TC) spectra were employed for each alloy. Using a mechanics-based data analysis, it was suggested that negative loads could be eliminated for the TC spectrum for low to intermediate maximum stress intensities. The suggestion was verified by subsequent testing. Using fractographic evidence, it was suggested that a further simplification in the spectra could be accomplished by eliminating low and intermediate peak load points resulting in near or below threshold maximum peak stress intensity values. It was concluded that load interactions become more important at higher stress intensities and with more plasticity at the crack tip. These results suggest that a combined mechanics/fractographic approach can be used to simplify other complex spectra.

Fatigue and fatigue crack growth characteristics of improperly quenched (substrength) aluminum

A program was conducted to determine how fatigue properties are affected in improperly quenched, substrength aluminum. Such a condition can result from an improper quench from the solutionizing temperature during heat-treating, leaving regions in the final product with substantially lower mechanical properties than those of properly heat-treated material. Properties thoroughly examined are both constant amplitude and spectrum loading fatigue, and fatigue crack growth, likewise under constant amplitude and spectrum loading conditions. Results obtained for two substrength aluminum alloys were similar: the improperly quenched material is inferior to properly quenched material under fatigue loading. For aluminum 2124, the fatigue crack propagation rate of the substrength material is over 40% faster than for the normal material, and only at stress intensities, approaching K c do the differences diminish. Also the effect of a defective quench appears to be influenced by alloy type, with aluminum 2124 affected to a much greater extent than 7075.

A review of some damage tolerance design approaches for aircraft structures

Damage tolerance design is becoming a necessity in the design of modern aircraft although its importance was recognized as long as four centuries ago by Leonardo da Vinci. Two decades ago structural design engineers and research workers felt the need of incorporating damage tolerance in the design of aircraft structure. Due to a lack of comprehensive damage tolerance methodology large scale component test results were used to develop empirical damage tolerance methods. Recently, linear elastic fracture mechanics has been used in predicting residual strength and crack growth rates in damaged structure. As a result of these efforts significant developments in cracked structure analytical methodology have been achieved. The recent Air Force requirement to apply linear elastic fracture mechanics approach in damage tolerance design of aircraft structures, warrants and critical review of various approaches. In this paper an attempt has been made to critically review some damage tolerance design approaches and their application to aircraft structures. The paper consists of three main sections: The first section reviews the residual strength analysis methodology, assumptions and limitations of each method are discussed through a simple example. The second part surveys the various crack propagation laws, including linear and non-linear ranges and spectrum loading effects. In the third and last section, fracture mechanics methodology is applied to several types of built-up structural components under spectrum loading conditions. The comparison of test results and analysis of complex structures indicate that simple methods of fracture mechanics can be applied to find the damage tolerant strength and rate of crack growth. The review presented in this paper indicates that the majority of work done in development of fracture mechanics analytical methodology has been based on data obtained from small scale laboratory specimens tested under closely controlled conditions of damage and environment. The validity of the methodology for complex structure under complex loading conditions has not been established. Before the results of a fracture mechanics analytical methodology can be accepted with a high degree of confidence many realism factors must be properly accounted for in the analysis.