Study of fatigue crack propagation in metallic structures

Tese de mestrado integrado. Engenharia Mecânica. Universidade do Porto. Faculdade de Engenharia. 201

Online fatigue crack growth monitoring with clip gauge and direct current potential drop

Fatigue is a well-known failure phenomenon which has been and still is extensively studied. Often structures are designed according to the safe-life principle so no crack initiation occurs. Nowadays there is a high emphasis on cost-efficiency, and one might rather opt for a fail-safe design. Therefore a certain amount of crack growth can be allowed in structures, but then a good knowledge of stresses and related crack growth rates is needed. To this end, extensive studies are done to obtain a material’s Paris law curve. Within the framework of research for offshore wind turbine constructions, tests were done to determine the crack growth rate of a high strength low alloy (HSLA) steel. A dedicated LabVIEW program was developed to be able to determine an entire Paris law curve with a single specimen, by controlling the stress intensity factor range (ΔK). The program is controlled by the readings of a clip gauge, which make it possible to plan the amount of crack growth per ΔK block and thus plan an entire test in advance. The potential drop technique was also applied in order to obtain the Paris law curve. Clip gauge results were compared with direct current potential drop monitoring. This comparison was done by means of an a/W-N diagram and the resulting Paris law curves. The results show a very good correlation between both methods and with the visual confirmation

Fatigue Crack Closure Analysis Using Digital Image Correlation

Fatigue crack closure during crack growth testing is analyzed in order to evaluate the critieria of ASTM Standard E647 for measurement of fatigue crack growth rates. Of specific concern is remote closure, which occurs away from the crack tip and is a product of the load history during crack-driving-force-reduction fatigue crack growth testing. Crack closure behavior is characterized using relative displacements determined from a series of high-magnification digital images acquired as the crack is loaded. Changes in the relative displacements of features on opposite sides of the crack are used to generate crack closure data as a function of crack wake position. For the results presented in this paper, remote closure did not affect fatigue crack growth rate measurements when ASTM Standard E647 was strictly followed and only became a problem when testing parameters (e.g., load shed rate, initial crack driving force, etc.) greatly exceeded the guidelines of the accepted standard

Assessing Variability in Microstructural Influence on Fatigue Crack Growth Behavior

The effect of microstructural variability has long been recognized as a major contributing factor in the scatter of published fatigue data. It is also acknowledged that these effects are generally more prevalent for short cracks and in the threshold region. A number of models exist to explain individual microstructural effects such as grain boundary influence, grain cluster, average grain size, porosity etc. It is the aim of the Aeronautics and Astronautics Fatigue Lab to develop an encompassing model that accurately predicts these effects. In order to develop this model a range of material data will be required to inform and validate the model simulation. It is the aim of this thesis to develop the methods required to generate suitable fatigue crack data and also image the crack propagation and strain fields. The methodology from ASTM E647 was used for the determination of crack growth data with the notable exception of the use of compression pre-cracking and relevant crack growth models for the ESE(T) specimen. Compression pre-cracking methods have been utilised as data have shown that standard pre-cracking methods may affect crack growth rate data and the determination of threshold values. High and low load ratio tests were conducted with closure accounted for, allowing for accurate determination of the fatigue crack growth threshold. High resolution DIC imagery was captured for a range of loads over a range of crack lengths and enabled the visualization of material strain fields. The imagery also allowed correlation between fatigue crack growth variability, closure data and the tortuosity of the crack surface

7075-T6 and 2024-T351 Aluminum Alloy Fatigue Crack Growth Rate Data

Experimental test procedures for the development of fatigue crack growth rate data has been standardized by the American Society for Testing and Materials. Over the past 30 years several gradual changes have been made to the standard without rigorous assessment of the affect these changes have on the precision or variability of the data generated. Therefore, the ASTM committee on fatigue crack growth has initiated an international round robin test program to assess the precision and variability of test results generated using the standard E647-00. Crack growth rate data presented in this report, in support of the ASTM roundrobin, shows excellent precision and repeatability

Temperature and load-ratio dependent fatigue-crack growth in the CrMnFeCoNi high-entropy alloy

Multiple-principal element alloys known as high-entropy alloys have rapidly been gaining attention for the vast variety of compositions and potential combinations of properties that remain to be explored. Of these alloys, one of the earliest, the ‘Cantor alloy’ CrMnFeCoNi, displays excellent damage-tolerance with tensile strengths of ∼1 GPa and fracture toughness values in excess of 200 MPa√m; moreover, these mechanical properties tend to further improve at cryogenic temperatures. However, few studies have explored its corresponding fatigue properties. Here we expand on our previous study to examine the mechanics and mechanisms of fatigue-crack propagation in the CrMnFeCoNi alloy (∼7 μm grain size), with emphasis on long-life, near-threshold fatigue behavior, specifically as a function of load ratio at temperatures between ambient and liquid-nitrogen temperatures (293 K–77 K). We find that ΔKth fatigue thresholds are decreased with increasing positive load ratios, R between 0.1 and 0.7, but are increased at decreasing temperature. These effects can be attributed to the role of roughness-induced crack closure, which was estimated using compliance measurements. Evidence of deformation twinning at the crack tip during fatigue-crack advance was not apparent at ambient temperatures but seen at higher stress intensities (ΔK ∼ 20 MPa√m) at 77 K by post mortem microstructural analysis for tests at R = 0.1 and particularly at 0.7. Overall, the fatigue behavior of this alloy was found to be superior, or at least comparable, to conventional cryogenic and TWIP steels such as 304 L or 316 L steels and Fe-Mn steels; these results coupled with the remarkable strength and fracture toughness of the Cantor alloy at low temperatures indicate significant promise for the utility of this material for applications at cryogenic environments

Crack growth from naturally occurring material discontinuities

All repairs to airframes now need to be assessed as to their effect on the damage tolerance the aircraft. To this end this chapter first discusses difference between the analysis tools needed for ab initio design and aircraft sustainment. It is shown that using small or physically short-crack da/dN versus Δ K data results in reduced through life costs and increased aircraft availability. The tests procedures needed to validate composite or supersonic particle deposition (SPD), repairs to operational aircraft are also discussed as is their relationship to the ASTM fatigue test standard E647-13a. This leads to an examination of the problem of crack growth from small naturally occurring material discontinuities under operational load spectra. A range of tools are available to account for crack growth in operational aircraft and several such tools are discussed, viz.: cycle-by-cycle analysis; the USAF characteristic K approach, etc. Specific attention is paid to the growth of lead cracks in operational aircraft which are shown to exhibit near exponential crack growth and to essentially have a cubic dependency on stress. It is shown that cracks growing in composite repairs exhibit the same crack length and stress dependency. This finding is then linked to current approaches which use the cubic rule to assess repairs to RAAF aircraft

Fatigue Crack Growth Tests and Analyses on a Ti-6Al-4V (STOA) Alloy using the Proposed ASTM Procedures for Threshold Testing

This thesis investigates fatigue crack growth rate behavior in the threshold and near-threshold regimes for a Ti-6Al-4V (STOA) alloy using two proposed ASTM procedures- (1) load-shedding (LS) using a larger load-shed rate than the current ASTM Standard E647 load-reduction (LR) test procedure, and (2) compression pre-cracking constant-amplitude (CPCA) or load-increasing (CPLI) and load-shedding (CPLS). Tests were conducted at a low stress ratio (R = 0.1) on compact C(T) specimens of two different widths (W = 51 and 76 mm) and threshold fatigue crack growth rates were generated. These test data were compared to previous test data produced from the same batch of material using the current LR and the CPCA test procedure. While no test procedure provided an exact representation of the threshold value (?Kth), the compression pre-cracking (CP) procedures were the most promising. The LR, LS, and CPLS test procedures were influenced by prior loading-history and various crack-closure mechanisms, leading to higher ?Kth values and slower crack growths in the threshold regime. The LS tests (at shed-rates of -0.08,-0.32, and -0.95 mm-1) generated ?Kth values that were 15% to 32% higher than the estimated threshold stress-intensity factor range (?*Kth)R=0.1. The CP test procedures are a more accurate alternative for developing near-threshold and threshold fatigue crack growth rates. The CPLS test procedure produced a ?Kth value that was 10% higher than (?*Kth)R=0.1. LR and LS tests produced different ?Kth values as a function of the specimen width for the given load ratio. The CP test procedures produced consistent crack growth rates over the same range of ?K values examined, independent of the specimen width. Further research is required for developing test procedure(s) capable of providing a more definitive representation of the ?Kth value and closureree fatigue crack growth rates in the threshold regime

On the USAF ‘risk of failure’ approach and its applicability to composite repairs to metal airframes

The USAF report on the risk analysis of aging aircraft fleets notes that the operational life of individual airframes is seldom equal to the design life of the fleet and that the life of an aircraft fleet t ends to be determined more by its inherent operational capability and maintenance costs rather than by the number of flight hours specified at the design stage. As such this paper focuses on whether the USAF approach to risk assessment can be used for airf rames repaired with a composite patch/doubler. To this end the present paper describes a test program designed to study the effect of adhesively -bonded composite repairs to fatigue cracks that, prior to repair, have grown from small naturally -occurring mat erials discontinuities. This study reveals that crack growth in composite repairs conforms to the exponential growth equation used in the USAF approach to assessing the risk of failure. Furthermore, the exponent, ω, in the exponential growth law can be de termined from the crack growth history associated with the unrepaired specimens and the simple reduction in the stress due to the application of the composite patch/doubler, using the ‘cubic rule’ that was previously used to assess crack growth in the RAAF F/A -18 (Hornet) fleet