Discrete crack growth analysis methodology for through cracks in pressurized fuselage structures

A methodology for simulating the growth of long through cracks in the skin of pressurized aircraft fuselage structures is described. Crack trajectories are allowed to be arbitrary and are computed as part of the simulation. The interaction between the mechanical loads acting on the superstructure and the local structural response near the crack tips is accounted for by employing a hierarchical modelling strategy. The structural response for each cracked configuration is obtained using a geometrically non-linear shell finite element analysis procedure. Four stress intensity factors, two for membrane behaviour and two for bending using Kirchhoff plate theory, are computed using an extension of the modified crack closure integral method. Crack trajectories are determined by applying the maximum tangential stress criterion. Crack growth results in localized mesh deletion, and the deletion regions are remeshed automatically using a newly developed all-quadrilateral meshing algorithm. The effectiveness of the methodology, and its applicability to performing practical analyses of realistic structures, is demonstrated by simulating curvilinear crack growth in a fuselage panel that is representative of a typical narrow-body aircraft. The predicted crack trajectory and fatigue life compare well with measurements of these same quantities from a full-scale pressurized panel test.     

Mixed mode fatigue growth of curved cracks emanating from fastener holes in aircraft lap joints

The finite element alternating method is extended further for analyzing multiple arbitrarily curved cracks in an isotropic plate under plane stress loading. The required analytical solution for an arbitrarily curved crack in an infinite isotropic plate is obtained by solving the integral equations formulated by Cheung and Chen (1987a, b). With the proposed method several example problems are solved in order to check the accuracy and efficiency of the method. Curved cracks emanating from loaded fastener holes, due to mixed mode fatigue crack growth, are also analyzed. Uniform far field plane stress loading on the plate and sinusoidally distributed pin loading on the fastener hole periphery are assumed to be applied. Small cracks emanating from fastener holes are assumed as initial cracks, and the subsequent fatigue crack growth behavior is examined until long arbitrarily curved cracks are formed near the fastener holes under mixed mode loading conditions.     

Reconstitution of fatigue crack growth in Al-alloy 2024-T3 open-hole specimens using microfractographic techniques

Fatigue is one of the main problems in the provision of service life and safety of aircraft structures. The menace of fatigue cracking is accentuated in areas of stress concentration, e.g., joints of structural components. An example is the fuselage where riveting is used. One of the techniques for improving the fatigue life of these connections is the cold expansion of the rivet hole. As part of a larger project on the fatigue behaviour of aeronautical structures, an experimental study of open-hole specimens in Al-alloy 2024-T3, with and without hole expansion, is presented. The residual stress field created by the cold expansion was experimentally assessed by using the X-ray technique and predicted by FEA. Fatigue tests were supplemented by SEM measurements of fatigue striation spacing along longitudinal and transverse directions in the crack surface of each specimen. Empirical models and fractographic techniques developed by Nedbal et al. are used for the analysis of the experimental data, and results of quantitative microfractography are presented. Crack tunnelling was quantified based on the reconstituted crack history and on the surface crack growth measurements.     

Experimental study of pressurized,foam reinforced,cracked cylindrical aluminum shells

Cracks in pressurized cylindrical shells are subject to stress fields peculiar to the shell geometry and loading conditions which makes them more critical than cracks in flat plates under similar conditions. In order to mitigate the stress field around the crack tips a foam layer is applied to the shell as reinforcement. The effect of this solution on crack behavior is evaluated experimentally. Pre-cracked cylindrical aluminum shells with layers of rigid foam applied to their inner side are tested by subjecting them to cyclic internal pressure. A test bench built especially for the purpose of cyclic testing is described as well as the testing procedure. Results indicate that crack growth rates are significantly reduced and fatigue life is extended by as much as 161%. The application of a foam layer to may help slow down the growth of existing cracks. This solution may prove useful especially in ageing aircraft fuselage structures.     

Effect of detail design on fatigue performance of fastener hole

In the present study, a series of tests were conducted on aluminum alloy 2024 to investigate the effect of detail design on the fatigue behavior of fastener holes in specimens. Two types of detail designs were concerned. One was the mode of fastener holes (countersunk rivet or countersunk bolt), the other was drilling process (traditional air-drilling process or one step compound cutting process). The fracture surfaces were observed by means of an optical microscope. Finite element method (FEM) was employed to analyze the distribution of stress around the fastener holes. The results showed that crack always initiated at the hole edge where the stress concentration occurred. Crack initiation was induced by stress concentration. Crack initiation life accounted for 80% of total fatigue life of fastener holes. The fatigue life of fastener hole using countersunk rivet was longer than that using countersunk bolt. Contrasted to traditional air-drilling process, the fatigue life of fastener hole could be improved by 44–55% using one step compound cutting process. However, the dispersibility of fatigue life became increasingly severe when fatigue life was prolonged.     

Practical use of the ‘equivalent’ measured stress intensity factor to control fatigue crack propagation rates in aircraft full-scale fatigue tests—First assessment of the method in testing of a pressurized aircraft fuselage

In the case of circumferential cracks in a cylindrical fuselage, the comparison of some analysis and test results shows that the theoretical stress intensity factor is a suitable correlation parameter of fatigue crack propagation rates, both in aircraft fuselages and in plane panels. Values of the ‘equivalent’ stress intensity factor, computed by applying the Barrois-Bhandari method to slot-opening measurements performed under decreasing loading levels, agree well with the values computed from two dimensional Theory of Elasticity, using the method of finite elements.

In the case of longitudinal cracks, the experimental values of the ‘equivalent’ stress intensity factor, i.e. the stress intensity factor of the infinite plane sheet containing a centre crack with the same elastic strain and stress distributions near the boundary of the plastically strained region around the crack tip, yield a good correlation of fatigue crack propagation rates of the cracked fuselage and of cracked plane structures. The values of the ‘equivalent’ stress intensity factor are lower than those of the theoretical stress intensity factor, the interest of which disappears, but are also far higher than the bidimensionally computed values, which are no longer to be considered.

Some meant of safety provided to limit crack openings will make it possible, in the near future, to investigate test conditions reaching the ultimate residual static strength of cracked structures, while avoiding, however, catastrophic failures.     

Static and high-rate loading of single and multi-bolt carbon–epoxy aircraft fuselage joints

Single-lap shear behaviour of carbon–epoxy composite bolted aircraft fuselage joints at quasi-static and dynamic (5 m/s and 10 m/s) loading speeds is studied experimentally. Single and multi-bolt joints with countersunk fasteners were tested. The initial joint failure mode was bearing, while final failure was either due to fastener pull-through or fastener fracture at a thread. Much less hole bearing damage, and hence energy absorption, occurred when the fastener(s) fractured at a thread, which occurred most frequently in thick joints and in quasi-static tests. Fastener failure thus requires special consideration in designing crashworthy fastened composite structures; if it can be delayed, energy absorption is greater. A correlation between energy absorption in multi-bolt and single-bolt joint tests indicates potential to downsize future test programmes. Tapering a thin fuselage panel layup to a thicker layup at the countersunk hole proved highly effective in achieving satisfactory joint strength and energy absorption.     

芯棒锥面结构对孔冷挤压强化残余应力场的影响

为了在紧固孔周引入均匀的残余压应力,以延长紧固孔构件的疲劳寿命、提高其抗应力腐蚀性能,利用ANSYS有限元软件,建立了轴对称弹塑性有限元模型,对直接芯棒冷挤压强化过程进行了仿真,特别是对芯棒的前锥段曲线结构形式进行了设计与分析,研究了前锥段曲线形式对残余应力场分布的影响.结果表明:孔壁表面的周向残余应力分布复杂且不均匀,比较而言,外凸型正弦曲线型芯棒所产生的残余压应力沿孔壁深度方向分布更加均匀;几种曲线形式的芯棒在上表面近孔边区域均产生了径向残余拉应力,在孔的挤入段产生了轴向残余拉应力,但外凸型正弦曲线型芯棒在上述区域所产生的残余拉应力较小,且分布区域也较小.     

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.     

Fatigue life prediction of interference fit fastener and cold worked holes

The purpose of this study is to present a methodology which estimates the fatigue life of interference fit fastener and cold worked holes. This methodology is mainly based on the determination of local stress with finite-element model computations and on the use of a multiaxial fatigue model. This approach is compared with experimental results carried out on test specimens representative of narrow body civil aircraft components. Analysis of the results shows a highly satisfactory correlation between prediction by calculation and experimental data.     

Residual strength analysis using CTOA criteria for fuselage structures containing multiple site damage

An extensive experimental program was conducted by the Boeing Company under the funding of the Federal Aviation Administration (FAA), National Aeronautics and Space Administration (NASA), and the United States Air Force Research Laboratory (USAF/RL) to investigate the effects of multiple-site damage (MSD) on the residual strength of typical fuselage splice joints. The experimental results were used to validate the analytical prediction using various methodologies, including STAGS (a generalized shell finite element code) with the crack-tip-opening angle and T^* fracture criteria.The test specimens consisted of large flat panels, curved panels, and an aft pressure bulkhead. The flat panel specimens included three types of longitudinal splice joints and one type of circumferential splice joint. For each type, one panel contained only a lead crack and the other two panels contained MSD 1.3 and 2.5 mm in size, respectively, at the fastener holes ahead of the lead crack. The curved panels were tested under simulated loads of combined cabin pressure and fuselage down bending. Two skin splice types were tested. For each splice type, one panel contained a lead crack only and the other had a lead crack with various sizes of MSD. A section of an aft fuselage containing a large lead crack and MSD in the pressure dome was also tested to demonstrate the capabilities of the methodologies in analyzing actual aircraft structures. This paper presents the analytical approaches and the comparison of predictions with the experimental results in terms of crack linkup stress and residual strength.     

Global behaviour of a composite stiffened panel in buckling. Part 2: Experimental investigation

The present study analyses an aircraft composite fuselage structure manufactured by the Liquid Resin Infusion (LRI) process and subjected to a compressive load. LRI is based on the moulding of high performance composite parts by infusing liquid resin on dry fibres instead of prepreg fabrics or Resin Transfer Moulding (RTM). Actual industrial projects face composite integrated structure issues as a number of structures (stiffeners, …) are more and more integrated onto the skins of aircraft fuselage.A post-buckling test of a composite fuselage representative panel is set up, from numerical results available in previous works. Two stereo Digital Image Correlation (DIC) systems are positioned on each side of the panel, that are aimed at correlating numerical and experimental out-of-plane displacements (corresponding to the skin local buckling displacements of the panel). First, the experimental approach and the test facility are presented. A post-mortem failure analysis is then performed with the help of Non-Destructive Techniques (NDT). X-ray Computed Tomography (CT) measurements and ultrasonic testing (US) techniques are able to explain the failure mechanisms that occured during this post-buckling test. Numerical results are validated by the experimental 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.     

力学在飞机结构设计中的应用与发展

首先阐述了飞机结构中的力学问题。基于我们近几年的研究工作,从若干方面介绍了力学在飞机结构设计中的应用与发展,如机身的广布疲劳损伤蔓延的断裂力学计算和风险评估,鸟撞驾驶舱盖和起落架落震的强度问题,以及低速冲击作用下机翼的复合材料层合板的脱层和基体断裂问题等。     

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.     

Global behaviour of a composite stiffened panel in buckling. Part 1: Numerical modelling

The present study analyses an aircraft composite fuselage structure manufactured by the Liquid Resin Infusion (LRI) process and subjected to a compressive load. LRI is based on the moulding of high performance composite parts by infusing liquid resin on dry fibres instead of prepreg fabrics or Resin Transfer Moulding (RTM). Actual industrial projects face composite integrated structure issues as a number of structures (stiffeners, …) are more and more integrated onto the skins of aircraft fuselage. A representative panel of a composite fuselage to be tested in buckling is studied numerically.     

Evaluation of the Effects of Corrosion on Fatigue Life of Clad Aluminum Alloy 2024-T3-Riveted Lap Joints with Acoustic Emission Monitoring

Corrosion affects the fatigue life of clad aluminum alloy-riveted lap joints, such as those found on an aircraft fuselage structure. Single-, double-, and triple-column-riveted lap joint specimens were fabricated and corroded in a Q-Fog accelerated corrosion chamber for five months using an ASTM G85-A5 prohesion test. Specimens were taken out of the chamber every 4?weeks, and the corrosion products which had been deposited on them were removed by immersion in concentrated nitric acid. For each corroded specimen, the mass loss with corresponding corrosion rate was determined. The specimens were fatigue loaded to failure on an MTS Universal Testing Machine with acoustic emission monitoring. Results indicate that exposure of lap joint specimens to this corrosive environment increased corrosion (mass loss), corrosion rate, and significantly reduced fatigue life. For a prolonged exposure in the corrosive environment, the fatigue life was reduced to zero, which has significant implication for aging aircraft. Acoustic emission monitoring successfully detected fatigue failure. Two failure modes, multisite crack damage and shear of the rivets, were observed.     

Observations and analyses of secondary bending for riveted lap joints

The phenomenon of secondary bending in riveted lap joints of the configuration representative of connections of the aircraft fuselage sheets in the longitudinal direction is investigated experimentally and analytically. The experiments involved strain gauge measurements of secondary bending stresses carried out in close proximity to the fatigue critical section of the riveted lap joint and fatigue tests performed to study the effect of secondary bending on the riveted joint fatigue life. The strain gauge measurement results allowed validation of a simple analytical model proposed by Schijve to estimate secondary bending moments induced in mechanically fastened joints with eccentricities. Variables considered in the fatigue tests were several joint geometry related parameters known to influence the magnitude of secondary bending in the fatigue critical location. It was shown that the fatigue test data for joints of various geometries, which were considerably scattered if the fatigue lives were presented against the applied stress amplitude, could become consolidated within common scatter bands when the lives were plotted in terms of the combined tensile stress amplitude including the bending stress computed from the model by Schijve.     

Effect of defects on fatigue life of plates with fastener holes

The behavior of defects (inclusions and cavities) on fatigue life of plates with fastener holes under tensile loading has been analyzed. Special attentions have been paid on the influence of the size and location of defects on fatigue life of plates with fastener holes. Thirty‐five different finite element models of plates with different size and location of defects are established and the nominal stress method is used to estimate the fatigue lives of this models. The results show that there is a region whose center is the maximum stress point of the hole without defects. When the defect is located in this region, the influence of defects on fatigue life of plates with fastener holes is obvious. When the defects are far away form this region' center, the defects hardly influence the fatigue life of plates with fastener hole. The larger the size of the defect is, the bigger this region is. In this region, the larger the size of the defect and the shorter the distance between the defect and the region's center, the shorter the fatigue life of plates with fastener hole is.     

On the finite element modeling of fatigue crack growth in pressurized cylindrical shells

A methodology for the computational modeling of the fatigue crack growth in pressurized shell structures, based on the finite element method and concepts of Linear Elastic Fracture Mechanics, is presented. This methodology is based on that developed by Potyondy [Potyondy D, Wawrzynek PA, Ingraffea, AR. Discrete crack growth analysis methodology for through crack in pressurized fuselage structures. Int J Numer Methods Eng 1995;38:1633–1644], which consists of using four stress intensity factors, computed from the modified crack integral method, to predict the fatigue propagation life as well as the crack trajectory, which is computed as part of the numerical simulation. Some issues not presented in the study of Potyondy are investigated herein such as the influence of the crack increment size and the number of nodes per element (4 or 9 nodes) on the simulation results by means of a fatigue crack propagation simulation of a Boeing 737 airplane fuselage. The results of this simulation are compared with experimental results and those obtained by Potyondy [1].