Fatigue crack growth in 7249‐T76511 aluminium alloy under constant‐amplitude and spectrum loading

Fatigue crack growth tests were conducted on compact, C(T), specimens made of 7249‐T76511 aluminium alloy. These tests were conducted to generate crack growth rate data from threshold to near fracture over a wide range of load ratios (R). Four methods were used to generate near threshold data: (1) ASTM E‐647 load reduction (LR), (2) compression pre‐cracking constant‐amplitude (CPCA), (3) compression pre‐cracking LR, and (4) constant crack mouth opening displacement LR method. A crack closure analysis was used to develop an effective stress‐intensity factor range against rate relation using a constraint factor (α = 1.85). Simulated aircraft wing spectrum tests were conducted on middle crack tension, M(T), specimens using a modified full‐scale fatigue test spectrum. The tests were used to develop the constraint‐loss regime (plane strain to plane stress; α = 1.85 to 1.15) behaviour. Comparisons were made between the spectrum tests and calculations made with the FASTRAN life prediction code; and the calculated crack growth lives were generally with ±10% of the test results.     

Flight-by-flight fatigue crack growth life assessment

Conventionally, fatigue crack growth in aircraft structures under flight spectrum loading is often analysed and predicted based on crack growth rates obtained from constant-amplitude crack growth testing with cycle-by-cycle life prediction methods or models. Because the mechanism of fatigue crack growth under spectrum loading is yet to be fully understood, no matter how closely the models are able to account for the load interaction effects, the predictions generally have to be subjected to the validation by fatigue crack growth tests using either representative specimens or real structures under the representative flight spectrum. In view of this fact, it is not difficult to deduce that the predictions should be much more reliable if the predictions are made directly based on the flight spectrum crack growth data. Therefore, a new approach to fatigue crack growth life assessment has been proposed in this paper based on the analysis of flight-by-flight fatigue crack growth data measured by quantitative fractography for several common aircraft structural materials under various fighter aircraft flight spectra. Quantitative fractography was successfully used for titanium coupons to generate crack growth curves under flight spectrum loading. The crack growths were also shown to be exponential. As a demonstration, the flight-by-flight approach was used to determine fatigue crack growth lives of aircraft aft fuselage frames under a fighter aircraft usage.     

Crack growth retardation under aircraft spectrum loads

In the past, the prediction of the fatigue life of a part was often compared with Miner's rule (Σn/N = 1) wherein damage was computed by a linear accumulation of growth under individual loads. The difference between the predicted and observed damage for variable amplitude tests showed up as either retarded or accelerated fatigue life depending on test conditions. In this program, fatigue crack growth tests were performed with the compact tension specimen under aircraft spectrum loading to develop techniques for predicting fatigue life. Two models were investigated for their application to retardation behavior—linear cumulative growth and Wheeler. Constant amplitude crack growth rate data represented by the Forman equation was used as source data for the computer models. These models appear suitable for comparison with actual test results for diffusion-bonded Ti-6A1-4V and HP-9Ni-4Co steel at 0.3 and 0.2 carbon contents for the aircraft spectrums investigated. The 2219-T851 aluminum alloy, on the other hand, exhibited accelelerated behavior in relation to the linear cumulative growth scheme. An explanation is proposed by relating recent crack closure and spike overload studies in aluminum with the format of the applied spectrum. Also, the Wheeler model was found dependent on material and spectrum variables, and the Forman equation did not accurately represent crack growth rate below a rates of 10⁻⁵ in/cycle.     

Block-by-block approaches for spectrum fatigue crack growth prediction

Two block-by-block approaches for improving spectrum fatigue crack growth prediction were proposed and developed in this paper from the observations and analyses of fatigue crack growth behaviours in either representative specimens or real aircraft structures under flight spectrum loading by using the quantitative fractography method. The first approach is the flight-by-flight approach that can be used to predict crack growth history curves for a tested spectrum crack growth data at different stress level for a critical location. The second approach called the effective block approach can be used to predict crack growth histories for un-tested spectra based on some previously tested spectrum crack growth data. In order to demonstrate the robustness of the block-by-block approaches for aircraft damage tolerance analysis, verification and consistency studies were conducted and presented using fatigue test results for different aircraft structures under several flight spectra. It was found that the block-by-block approaches are able to provide significant advantages over conventional fatigue lifing approaches for aircraft damage tolerance analysis.     

TA1异质自冲铆接头力学性能及失效机理

通过自冲铆接对比试验获得接头最优铆接参数,并以此制备TA1钛合金板分别与Al5052铝合金板和H62铜合金板的异质自冲铆接头。通过静力学实验和疲劳实验研究异质接头的力学性能,并运用疲劳三参数经验公式拟合S-N曲线,最后利用扫描电镜和能谱仪进行断口分析和能谱分析进而研究接头的疲劳失效机理。结果表明,TA1-H62(STH)接头静失效载荷优于TA1-Al5052(STA)接头;且前者在低载荷下疲劳寿命优于后者,STA接头则在高载荷下优势明显。STA接头疲劳失效模式为下板断裂,STH接头则出现了两种失效模式;两板间及铆钉与上下板之间接触区域发生的剧烈微震磨损是导致疲劳裂纹萌生的主要原因。     

Retardation effects due to overloads in aluminium‐alloy aeronautical components

Fatigue data are generally derived under constant‐amplitude loading conditions, but aircraft components are subjected to variable‐amplitude loading. Without interaction effects, caused by overloads and underloads intermingled in a loading sequence, it could be relatively easy to establish a crack growth curve by means of a cycle‐by‐cycle integration. However, load‐spectrum effects largely complicate a crack growth under variable‐amplitude cycling. In this paper, fatigue crack growth behaviour of aeronautical aluminium alloy 2024‐T3 was studied. Effects of various loading conditions such as stress ratio and amplitude loadings were investigated. In particular, the effect of different overloads on the fatigue crack growth was simulated using Zencrack code. Preliminary analyses on Compact Tension (CT) specimens proved that the numerical results generated were in agreement with the results provided by an afgrow code for the same conditions. A case study was carried out on a helicopter component, undergoing repeated overloads, to compare numerical results obtained implementing yield zone models in Zencrack.     

Fatigue crack growth and delamination mechanisms of Ti/CFRP fibre metal laminates at high temperatures

Ti/CFRP (titanium/carbon fibre reinforced polymer) fibre metal laminates (FMLs) are composed of titanium sheets and carbon fibres reinforced PMR (polymerization of monomeric reactants) type polyimide resin. Due to the outstanding heat resistance of the material, it can be used in hypersonic aircraft applications. Fatigue cracks in the metal layer and delamination at metal/fibre interface may occur in long‐term high‐temperature use processes. However, the behaviour of the fatigue failure at high temperatures has not been investigated. A temperature‐dependent equation has not been presented to predict the crack growth behaviour at high temperatures. In this study, to investigate the crack propagation and delamination behaviours, fatigue crack growth rate tests using tension‐tension loads at 25°C, 80°C, 120°C, and 150°C were conducted in accordance with ASTM E647‐15e1. The results indicated that the variation in fatigue crack growth rate could be described by a modified temperature‐dependent Paris equation. Interfacial strength and tensile strength may influence fatigue failure at high temperatures. Hence, these strength values were also obtained to analyse the mechanism of fatigue behaviour. The delamination area increased exponentially with temperature due to the weakening of the Ti/CFRP interface, and delamination was invariably generated on the microcracks of the titanium layers.     

Improved test method and analytical modelling for fatigue crack growth in coarse‐grain titanium alloy with rough fatigue surfaces

Fatigue crack growth rate properties are typically determined by experimental methods in accordance with ASTM Standard E647. These traditional methods use standard notched specimens that are precracked under cyclic tensile loads before the main test. The data that are produced using this approach have been demonstrated elsewhere to be potentially adversely affected by the test method, particularly in the threshold region where load reduction (LR) methods are also required. Coarse‐grained materials that exhibit rough and tortuous fatigue surfaces have been observed to be strongly affected by the tensile precracking and LR, in part because the anomalies caused by crack closure and roughness‐induced closure become more important. The focus of the work reported in this paper was to further develop methods to determine more accurate fatigue crack growth rate properties from threshold through to fracture for coarse‐grained, β‐annealed, titanium alloy Ti‐6Al‐4V extra low interstitial thick plate material. A particular emphasis was put upon the threshold and near threshold region, which is of strong importance in the overall fatigue life of components. New approaches that differ from the ASTM Standard included compression precracking, LR starting from a lower load level and continuing the test beyond rates where crack growth would otherwise be considered below threshold. For the threshold regime, two LR methods were also investigated: the ASTM method and a method where the load is reduced with crack growth such that the crack mouth opening displacement is held constant, in an attempt to avoid remote closure. Constant amplitude fatigue crack growth rate data were produced from threshold to fracture for the titanium alloy at a variety of stress ratios. Spike overload tests were also conducted These data were then used to develop an improved analytical model to predict crack growth under spectrum loading and the predictions were found to correlate well with test results.     

Fatigue crack propagation under in‐phase and out‐of‐phase biaxial loading

Fatigue damage characteristics of aluminium alloy under complex biaxial loads such as in‐phase and out‐of‐phase loading conditions and different biaxiality ratios have been investigated. The effects of microscale phenomena on macroscale crack growth were studied to develop an in‐depth understanding of crack nucleation and growth. Material characterization was conducted to study the microstructure variability. Scanning electron microscopy was used to identify the second phase particles, and energy dispersive X‐ray spectroscopy was performed to analyse their phases and elements. Extensive quasi‐static and fatigue tests were conducted on Al7075‐T651 cruciform specimens over a wide range of load ratios and phases. Detailed fractography analysis was conducted to understand the crack growth behaviour observed during the fatigue tests. Significant differences in crack initiation and propagation behaviour were observed when a phase difference was applied. Primarily, crack retardation and splitting were observed because of the constantly varying mode mixity caused by phase difference. The crack growth behaviour and fatigue lives under out‐of‐phase loading were compared with those under in‐phase loading to understand the effect of mixed‐mode fracture.     

The effects of load sequencing on the fatigue life of 2024-T3 aluminum alloy

Current fatigue life analysis of metallic rotorcraft dynamic components are based on a linear cumulative damage concept known as Miner's rule. This type of analysis assumes that there is no load sequence effect that occurs during the fatigue loading history. Past studies have shown that linear cumulative damage analysis of fatigue tests has produced life predictions that have been conservative as well as unconservative. Some of this uncertainty has been attributed to the fact that load sequence effects do exist in most fatigue load spectra. As a first phase the study reported herein was done to evaluate the load sequencing effects that could exist in commercial fixed-wing fatigue load spectra. To evaluate these effects a typical commercial wing spectra was reordered using a scheme that had previously been shown in fatigue block loading to produce the shortest fatigue lives. This scheme starts with the smallest load range in a load sequence and proceeds in ascending order until the largest load range is reached. Tests on open hole test specimens made of 2024-T3 aluminum alloy were conducted on the normal sequence of loads as well as the reordered scheme called lo–hi. Test results showed no significant differences between the fatigue lives of the normal load sequence and the reordered load sequence. A computer program called FASTRAN which calculates total fatigue life using only crack growth data was shown to predict the fatigue life of the spectrum tests with acceptable accuracy.     

FATIGUE OF THICK-SECTION COLD-EXPANDED HOLES WITH AND WITHOUT CRACKS

Abstract— Fatigue tests under spectrum loading were conducted to evaluate hole cold-expansion in thick-section open-hole aluminium alloy specimens, some of which contained residual fatigue cracks before cold expansion. Cold expansion resulted in an increase in life by a factor of about 7. Small residual fatigue cracks did not inhibit the effectiveness of the cold-expansion process, indicating that it may not be essential to remove such cracks prior to hole cold-expansion. The increase in life is primarily associated with a reduced crack propagation rate compared with that for cracks from non-cold-expanded holes. Cold-expanded hole fractures displayed a marked disparity in crack depths adjacent to the two faces of the specimens. Considerable differences were evident in crack depths and fatigue crack areas at failure between cold-expanded and non-cold-expanded hole specimens. These findings have ramifications in the damage tolerance assessment of aircraft structures.     

Crack growth predictions using a strip‐yield model for variable‐amplitude and spectrum loading

Fatigue‐crack‐growth tests were conducted on compact, C(T), specimens made of D16Cz aluminum alloy. Constant‐amplitude tests were conducted over a range of stress ratios (R = P_min/P_max = 0.1 to 0.75). Comparisons were made between test data from middle‐crack tension, M(T), specimens from the literature and C(T) specimens. A crack‐closure analysis was used to collapse the rate data from both specimen types into a fairly narrow band over many orders of magnitude in rates using proper constraint factors. Constraint factors were established from single‐spike overload and constant‐amplitude tests. The life‐prediction code, FASTRAN, which is based on the strip‐yield‐model concept, was used to calculate the crack‐length‐against‐cycles under constant‐amplitude (CA) loading and the single‐spike overload (OL) tests; and to predict crack growth under variable‐amplitude (VA) loading on M(T) specimens and simulated aircraft loading spectrum tests on both specimen types. The calculated crack‐growth lives under CA and the OL tests were generally within ±20 % of the test results, the predicted crack‐growth lives for the VA and Mini‐Falstaff tests on the M(T) specimens were short by 30 to 45 %, while the Mini‐Falstaff+ results on the C(T) specimens were within 10 %. Issues on the crack‐starter notch effects under spectrum loading are discussed, and recommendations are suggested on avoiding these notch effects.     

Effect of salt-water fog on fatigue crack growth behaviour of 7050 aluminium alloy in different orientations

The fatigue crack growth behaviour of 7050 T73651 high strength aluminium alloy that was originally developed for the aircraft industry was investigated in this study. The tests were conducted by using C-T specimens machined in six orientations under the action of constant amplitude sinusoidal load cycles. The tests were first carried out in laboratory air and then repeated in salt-water fog of a 5% NaCl solution to observe the effect of the environment on the fatigue crack growth behaviour. The experimental results showed that the fatigue life, maximum stress intensity range and the fatigue crack growth rate of the specimens were seriously affected by the environment. The severity of the effect, on the other hand, was observed to be dependent on the orientation. The strongest orientation was determined to be L-S, while the weakest was S-L.     

Fatigue behaviour and life prediction of fibre reinforced metal laminates under constant and variable amplitude loading

ABSTRACT Fatigue crack growth of fibre reinforced metal laminates (FRMLs) under constant and variable amplitude loading was studied through analysis and experiments. The distribution of the bridging stress along the crackline in centre‐cracked tension (CCT) specimen of FRMLs was modelled numerically, and the main factors affecting the bridging stress were identified. A test method for determining the delamination growth rates in a modified double cracked lap shear (DCLS) specimen was presented. Two models, one being fatigue‐mechanism‐based and the other phenomenological, were developed for predicting the fatigue life under constant amplitude loading. The fatigue behaviour, including crack growth and delamination growth, of glass fibre reinforced aluminium laminates (GLARE) under constant amplitude loading following a single overload was investigated experimentally, and the mechanisms for the effect of a single overload on the crack growth rates and the delamination growth rates were identified. An equivalent closure model for predicting crack‐growth in FRMLs under variable amplitude loading and spectrum loading was presented. All the models presented in this paper were verified by applying to GLARE under constant amplitude loading and Mini‐transport aircraft wing structures (TWIST) load sequence. The predicted crack growth rates are in good agreement with test results.     

Bonded aircraft repairs under variable amplitude fatigue loading and at low temperatures

Bonded repairs can replace mechanically fastened repairs for aircraft structures. Compared to mechanical fastening, adhesive bonding provides a more uniform and efficient load transfer into the patch, and can reduce the risk of high stress concentrations caused by additional fastener holes necessary for riveted repairs. Previous fatigue tests on bonded Glare (glass‐reinforced aluminium laminate) repairs were performed at room temperature and under constant amplitude fatigue loading. However, the realistic operating temperature of ?40 °C may degrade the material and will cause unfavourable thermal stresses. Bonded repair specimens were tested at ?40 °C and other specimens were tested at room temperature after subjecting them to temperature cycles. Also, tests were performed with a realistic C‐5A Galaxy fuselage fatigue spectrum at room temperature. The behaviour of Glare repair patches was compared with boron/epoxy ones with equal extensional stiffness. The thermal cycles before fatigue cycling did not degrade the repair. A constant temperature of ?40 °C during the mechanical fatigue load had a favourable effect on the fatigue crack growth rate. Glare repair patches showed lower crack growth rates than boron/epoxy repairs. Finite element analyses revealed that the higher crack growth rates for boron/epoxy repairs are caused by the higher thermal stresses induced by the curing of the adhesive. The fatigue crack growth rate under spectrum loading could be accurately predicted with stress intensity factors calculated by finite element modelling and cycle‐by‐cycle integration that neglected interaction effects of the different stress amplitudes, which is possible because stress intensities at the crack tip under the repair patch remain small. For an accurate prediction it was necessary to use an effective stress intensity factor that is a function of the stress ratio at the crack tip R_{crack tip} including the thermal stress under the bonded patch.     

Crack growth predictions in a complex helicopter component under spectrum loading

Fatigue crack growth predictions have been made on a helicopter round‐robin crack configuration. The crack configuration was a small corner defect at the edge of a large central hole in a flanged plate made of 7010 aluminium alloy and the component was subjected to a simulated helicopter spectrum loading. The crack growth rate data and the stress‐intensity factor (K) solution for the crack configuration were provided in the round‐robin. The FASTRAN life‐prediction code was used to predict fatigue crack growth under various load histories on the aluminium alloy, such as Rotorix and Asterix, on both compact tension C(T) specimens and the complex crack configuration. A BEASY three‐dimensional stress‐intensity factor solution for the round‐robin problem was also provided for this paper and is compared with the original K solution. Comparisons are made between measured and predicted fatigue crack growth lives for both crack configurations. The predicted lives for the C(T) specimens were 15–30% longer than the measured lives; and crack growth in the round‐robin configuration agreed very well in the early stages of crack growth, but the life was 30% short of the test results at the final crack length.     

Experiments and model predictions for fatigue crack propagation in riveted lap‐joints with multiple site damage

In this paper, the growth of long fatigue cracks up to failure in aircraft components is studied. A deterministic model is presented, able to simulate the growth of fatigue through cracks located at rivet holes in lap‐joint panels. It also includes criteria to assess the link‐up of collinear adjacent cracks in a MSD scenario. To validate the model, a fatigue test campaign was carried out on riveted lap‐joint specimens in order to produce experimental crack growth and link‐up data. Accurate measurements of naturally occurred surface cracks were automatically performed by the Image Analysis technique, thus allowing the tests to run 24 h a day. The comparison between experimental tests and numerical simulations is good, thus confirming the model as a useful tool for the assessment of fatigue life of aircraft riveted joints.     

Fatigue life time prediction of PoAF Epsilon TB-30 aircraft – Implementation of automatic crack growth based on 3D finite element method

The objective of this work is to predict the fatigue lifetime (TVF) of the Portuguese Air Force (PoAF) Epsilon aircraft based on the computational fatigue crack growth modelling.The spectra of loads were used in experimental tests of two specimen series (designed to simulate in the laboratory the critical area of the aircraft) to assess experimentally the difference between the PoAF and manufacture spectra.In order to predict the TVF by a generic spectrum was computationally implemented a methodology for automatic crack propagation. Through the development of a interface between ANSYS and MATLAB was possible to determine the stress intensity factors and hence the geometric factor for the specimen geometry which was designed by PoAF in previous works. The stress intensity factors were validated with the methods available in the literature: Pickard, Pommier and Newman.The spectra of charges and the geometric factor allowed the computer implementation of the following propagation laws: Paris, NASGRO, Walker, Forman and Wheeler.Finally, it was established for the PoAF operation the new inspections plans according to the manufacturer methodology by making an extrapolation of real scale test results obtained with the manufacturer spectrum.At the end of the article the authors concluded that the TVF of Epsilon aircraft is 24,500 flight hours (FH), the first inspection should be done when the aircraft reaches 10,000 FH and the flowing inspections should be done with a periodicity of 3000 FH until the crack reaches a critical dimension of 1.5 mm.     

Small-crack growth and fatigue life predictions for high-strength aluminium alloys. Part II: crack closure and fatigue analyses

Small-crack effects were investigated in two high-strength aluminium alloys: 7075-T6 bare and LC9cs clad aluminium alloys. Both experimental and analytical investigations were conducted to study crack initiation and growth of small cracks. In the experimental program, fatigue and small-crack tests were conducted on single-edge-notch tension (SENT) specimens and large-crack tests were conducted on middle-crack tension specimens under constant-amplitude and Mini-TWIST spectrum loading. 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 equations for surface and corner cracks at the notch in the SENT specimen. (Part I was on the experimental and fracture mechanics analyses and was published in Fatigue Fract. Engng Mater. Struct. 21 , 1289–1306, 1998.) This part focuses on a crack closure and fatigue analysis of the data presented in Part I. A plasticity-induced crack-closure model was used to correlate large-crack growth rate data to develop the baseline effective stress intensity factor range (Δ K _eff ) against rate relations for each material, ignoring the large-crack threshold. The model was then used with the Δ K _eff rate relation and the stress intensity factors for surface or corner cracks to make fatigue life predictions. The initial defect sizes chosen in the fatigue analyses were similar to those that initiated failure in the specimens. Predicted small-crack growth rates and fatigue lives agreed well with experiments.