Residual strength of unstiffened aluminum panels with multiple site damage

This study investigated the residual strength of unstiffened aluminum panels with widths of 381 mm and 229 mm containing multiple site damage (MSD). The MSD usually occurs at rivet holes, or other stress concentration locations within an aircraft structure. This study simulated rivet holes with MSD, by using holes of constant diameter with small cracks, evenly spaced across the midspan of specimens. The panels were prepared by either fatiguing MSD damage at rivet holes or simulating fatigue damage by saw cuts at each hole. Each specimen was subjected to a monotonically increasing tensile load until failure occurred across the midspan of the gauge section. Five different failure criteria which do not model the stable crack extension were evaluated to predict the residual strength (failure load) for each specimen geometry. These criteria provided a wide range of residual strength predictions for wide and narrow panels with MSD. A failure criterion which involved the plastic zone (yielding) of the lead and MSD cracks gave the most accurate prediction of failure load for panels with MSD damage. The width of the specimens did not affect, in general, the trends in the prediction of failure loads from the five failure criteria.     

Foreign object damage and fatigue crack threshold: Cracking outside shallow indents

Foreign Object Damage (FOD) usually happens when objects are ingested into jet engines powering military or civil aircraft. Under extreme conditions, FOD can lead to severe structural damage. More commonly it produces local impacted sites of the fan and compressor airfoils, lowering fatigue life of these components. FOD is a prime cause for maintenance and repair in aircraft engines. In this paper, a framework for analyzing FOD and its effect on fatigue cracking is established. A finite element analysis is used to identify three relevant regimes of FOD related to the depth of penetration into the substrate, and to determine the residual stresses. Most of the emphasis in this paper focuses on fatigue cracks emerging from shallow indentations, which are generally expected to be of most practical concern. Full three-dimensional finite element solutions are obtained for semi-circular surface cracks emerging from specific locations at the indentation revealing the influence of the residual stress on the stress intensity factor distribution. For shallow indents, a relatively simple dimensionless formula for the relation between the residual stress intensity factor, the crack size, and the indentation width are developed. These results, together with results for the intensity factor variations due to cyclic loading, have been used to address the question: To what extent do the residual stresses caused by the FOD reduce the critical crack size associated with threshold fatigue crack growth? Formulas for the critical crack size are obtained. Specific results are presented for the blade alloy, Ti-6Al-4V, revealing that FOD can reduce the critical crack size by as much as 60%.     

MULTIPLE-SITE DAMAGE IN AIRCRAFT FUSELAGE STRUCTURES

Abstract— Since the Aloha accident the multiple site damage (MSD) problem of riveted lap joints in aircraft fuselages has drawn much attention. The failure scenarios for a lead crack and more small MSD cracks as discussed by Broek and Swift are summarized, including recent results of a relevant test series by Broek. It shows that small MSD cracks can significantly reduce the load for unstable crack extension. Prevention of catastrophic consequences requires crack arresting capability of the structure. Related aspects of the problem are discussed with reference to failure criteria for ligament failure, the MSD problem for existing and new aircraft, and different options for crack stopper bands.     

Fracture analysis of multi-site cracking in fuselage lap joints

A two-dimensional plane stress elastic fracture mechanics analysis of a cracked lap joint fastened by rigid pins is presented and results are applied to the problem of multi-site damage (MSD) in riveted lap joints of aircraft fuselage skins. Two problems are addressed, the problem of equal length MSD cracks and the problem of alternating length MSD cracks. For the problem of equal length cracks, two models of rivet/skin interactions are studied and the role of residual stresses due to the riveting process is explored. Stress intensity factors are obtained as a function of normalized crack length. Also, the load distribution among rivet rows and the compliance change of the joint due to MSD cracking are obtained. For the problem of alternating length cracks, attention is focussed on how load is distributed between columns of rivets and how this load shedding can alter crack tip stress intensity factors. The equal and alternating length crack analyses reveal no clear-cut mechanism to explain the relative uniformity of fatigue cracks emerging from lap joint rivet holes in actual aircraft and in mechanical lap joint tests.     

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.     

Compliance measurements for assessing structural integrity

The phenomenon of aging structures has focused attention on the problems of multiple-site damage (MSD) and widespread fatigue damage (WFD). In Australia, the problem was highlighted by the November 1990 failure of a Royal Australian Air Force (RAAF) Macchi aircraft which suffered a port wing failure whilst in an estimated 6 g maneuvre and by the September 1998 explosion at the EXXON gas plant in Victoria. To assist in the understanding and the management of this problem the present paper uses the newly developed finite element alternating technique, for an arbitrary number of interacting three dimensional cracks, which we refer to as the MSD FEAT algorithm, to evaluate whether compliance measurements are useful in assessing continuing airworthiness. Traditionally the MSD FEAT and the FEAT analysis tools, i.e. the analysis methodology for a single crack, have been used only to analyse the stress intensity factor distributions around crack faces. The new work described in this paper enables the displacement field, and hence the compliance, to be calculated at any given location within the structure. Initial results confirm that this technique produces correct displacements and is capable of determining the crack tip opening displacement to within ∼0.7% for semi-elliptical surface flaws. Earlier work conducted on two dimensional MSD problems found that when using compliance measurements to evaluate cracking there was an optimal sensor length for monitoring crack interaction effects. The present paper extends this study to three-dimensional flaws via a combined analytical and experimental research program. The experimental work focuses on specimens containing two interacting quarter elliptical cracks. Here the changes in compliance of the specimen under monotonic loading, and fracture load, were measured and were found to be in good agreement with those predicted using the newly developed MSD FEAT algorithm. Results of this analysis indicate that placement of sensors in an optimal position is crucial.     

Residual strength of aircraft panels with multiple site damage

Multiple site damage (MSD) is a type of cracking that may be found in aging airplanes which can adversely affect the damage tolerance of an airframe structure. In this paper the behavior of MSD is studied by examining the interaction of cracks in stiffened and riveted panels. The hybrid finite element method, in conjunction with the complex variable theory of elasticity, is used to provide accurate and efficient solutions to these problems. Typical results include stress intensity factors at the crack tips, stress concentration factors in the stiffeners, and rivet loads for a stiffened structure with multiple cracks. Particular emphasis is placed on the derivation and interpretation of residual strength diagrams. This study produces a better understanding of the interaction between multiple cracks and provides insight for avoiding MSD in future designs.This work was supported by the Technical Center, Federal Aviation Administration, U.S. Department of Transportation     

Influence of MSD crack pattern on the residual strength of flat stiffened sheets

A parameter study of the residual strength for a multiple site damaged (MSD) stiffened sheet is presented. The analysis is based on an elastic-plastic fracture analysis using the yield-strip model for interaction between a lead crack and the smaller MSD cracks. Two crack growth criteria, one with a pronounced crack growth resistance and one with no crack growth resistance and five different MSD crack patterns, are analysed for different sizes of the lead crack and the smaller MSD cracks. The analysis indicates that the residual strength reduction depends on all these parameters and that MSD may totally erode the crack arrest capability of a tear strap. Another important outcome is that for certain combinations also very small MSD cracks may induce a significant residual strength reduction.     

The residual strength characteristics of stiffened panels containing fatigue cracks

In heavy structural members, where plane strain conditions prevail, linear fracture mechanics can be used for predicting residual strength. Aircraft structures consist largely of sheet structures with plane stress conditions where linear fracture mechanics do not seem to apply. Yet it is in the aircraft main structure that large fatigue cracks can develop and that has to be designed fail-safe. The present paper describes a method to predict the residual strength of a cracked sheet structure.Contrary to an unstiffened sheet, the sheet structure contains stiffening elements that can act as crack stoppers. This crack arresting action and its consequences for the residual strength are considered in the analysis.The paper proposes a method that relates the crack resistance of a stiffened panel to that of an unstiffened sheet. It takes full account of sheet-stringer interaction in the cracked region. A criterion for crack arrest is put forward. Ultimate panel failure after crack arrest is initiated either by subsequent unstable crack growth or by stiffener failure. Critical load conditions for both failure modes are presented. In case crack arrest does not occur, the residual strength of the unstiffened panel constitutes a safe lower bound.Computational results of the interacting rivet forces by both analytical and numerical (finite element) methods are presented. From these the load concentration in the stiffener and the reduction of the stress intensity at the crack tip can be determined. This enables the complete residual strength characteristics to be predicted.The results of residual strength tests on bonded and riveted panels with symmetric strip stiffeners or eccentric Z-stringers fully substantiate the method proposed for residual strength calculations.     

Analysis of fracture and delamination in laminates using 3D numerical modelling

The static failure behaviour of the fibre-metal laminate GLARE is examined using 3D finite element simulations. The configuration analysed is a centre-cracked tensile specimen composed of two aluminium layers sandwiching a cross-plied, fibre-epoxy layer. The crack and delamination growths are simulated by means of interface elements equipped with a mixed-mode damage model. The mode-mixity is derived from an energy criterion typically used in linear elastic fracture mechanics studies. The damage kinetic law is rate-dependent, in order to simulate rate effects during interfacial delamination and to avoid numerical convergence problems due to crack bifurcations. The numerical implementation of the interface damage model is based on a backward Euler approach. In the boundary value problem studied, the failure responses of GLARE specimens containing elastic aluminium layers and elasto-plastic aluminium layers are compared. The development of plastic deformations in the aluminium layers stabilizes the effective failure response, and increases the residual strength of the laminate. For a ‘quasi-brittle’ GLARE specimen with elastic aluminium layers, the residual strength is governed by the toughness for interfacial delamination, and is in close correspondence with the residual strength obtained from a closed-form expression derived from energy considerations. Conversely, for a ‘ductile’ GLARE specimen with elasto-plastic aluminium layers, the residual strength is also determined by the relation between the fracture strength and the yield strength of the aluminium. The amount of constraint by the horizontal displacements at the vertical specimen edges has a moderate to small influence on the residual strength. Furthermore, the ultimate laminate strength is lower for a larger initial crack length, and shows to be in good correspondence with experimental values.     

有限板共线多孔MSD疲劳裂纹扩展有限元模拟

介绍了二维断裂分析有限元软件FRANC2D/L在疲劳裂纹扩展模拟方面的基本步骤．利用该软件对有限板中心孔边对称裂纹的疲劳裂纹扩展进行模拟计算,对比模拟结果和试验数据,发现两者吻合良好,证明了利用该方法模拟疲劳裂纹扩展的可靠性．将FRANC2D/L应用到有限板共线多孔MSD疲劳裂纹扩展的有限元模拟上,得到了各孔边裂纹的长度和疲劳扩展寿命之间的关系曲线．模拟计算结果表明,在相同条件下,有限板中心孔边对称裂纹的裂纹扩展寿命要远远高于MSD结构中中心孔边裂纹的疲劳扩展寿命;由于MSD结构中影响各孔边裂纹的因素有所差异,各条裂纹的疲劳扩展寿命也会有所不同．另外,还给出了不含主裂纹的MSD和含主裂纹的MSD两种情况下的疲劳裂纹扩展历程,通过比较得知,含主裂纹的MSD结构更容易发生裂纹的合并和贯穿致使结构发生破坏．     

Onset of multiple site damage and widespread fatigue damage in aging airplanes

This paper presents a methodology for predicting the thresholds of multiple site damage and widespread fatigue damage in fuselage lap slices. Widespread fatigue damage is a type of multiple cracking that reduces the airframe residual strength to a level below the damage tolerant requirement. The MSD threshold refers to the point in the lifetime of an airplane when two adjacent collinear fatigue cracks can linkup at the allowable stress. The WFD threshold is the point in time when linkup of a primary crack created from accidental damage and secondary cracks created from fatigue can result in a catastrophic failure. The methodology presented in this paper combines results from residual strength analysis and fatigue crack growth testing to determine these thresholds. In particular, a displacement compatibility approach is adopted to calculate residual strength in curved stiffened panels tested in the laboratory. The laboratory experiments also include fatigue testing of full-scale panels containing a debonded lap slice. Based on this methodology, the threshold of widespread fatigue damage for these laboratory panels, adjusted to zero minimum-stress cycling, is between 32,000 and 40,000 cycles, and the threshold of multiple site damage is about 70,000 cycles. This revised version was published online in August 2006 with corrections to the Cover Date.     

Ensuring structural damage tolerance of russian aircraft

This paper considers most common problems related with structural integrity of civil aircraft in Russia taking into account the development of regulatory requirements, prevention of multiple site fatigue damages, improvements of crack resistance of structural materials, optimizations of aircraft type structures, development of methods for residual strength analyses of stiffened structures as well as for crack growth rates under random service loading spectra, experimental results for crack resistance degradation, methods to prevent structural failure for long operated aircraft due to corrosion.     

Simulation of fatigue crack growth in components with random defects

The paper presents a probabilistic method for the simulation of fatigue crack growth from crack-like defects in the combined operating and residual stress fields of an arbitrary component. The component geometry and stress distribution are taken from a standard finite element stress analysis. Number, size and location of crack-like defects are ‘drawn’ from probability distributions. The presented fatigue assessment methodology has been implemented in a newly developed finite-element post-processor, P • FAT, and is useful for the reliability assessment of fatigue critical components. General features of the finite element post-processor have been presented. Important features, such as (i) the determination of the life-controlling defect, (ii) growth of short and long cracks, (iii) fatigue strength and fatigue life distribution and (iv) probability of component fatigue failure, have been treated and discussed. Short and long crack growth measurements have been presented and used for verification of the crack growth model presented.     

Fatigue growth of multiple-cracks near a row of fastener-holes in a fuselage lap-joint

The fatigue growth of multiple cracks, of arbitrary lengths, emanating from a row of fastener holes in a bonded, riveted, lap joint in a pressurized aircraft fuselage is studied. The effects of residual stresses due to a rivet misfit, and of plastic deformation near the hole, are included. A Schwartz-Neumann alternating method which uses the analytical solution for a row of multiple colinear cracks in an infinite sheet (the crack-faces being subject to arbitrary tractions), is developed to analyze this MSD problem on a personal computer. It is found that for a range of crack lengths, a phenomena wherein the shorter cracks may grow faster than longer cracks may exist.     

Application of fracture mechanics to aircraft structural safety

Design requirements have been developed to insure structural safety for current and future USAF aircraft. New structure is assumed to be flawed. Materials, stress levels and structural arrangements are chosen so as to prevent damage from growing to catastrophic size prior to detection. Safety from damage induced through service usage is insured by providing inspection capability and by meeting specific residual strength and safe crack growth requirements. Compliance with these requirements implies the capability to predict growth rates under complex loadings and to calculate the fracture strength of structures fabricated of relatively tough materials which may exhibit large amounts of crack tip plasticity prior to failure. This paper will review significant factors leading to the development of damage tolerance criteria and illustrate the role of fracture mechanics in the analysis and testing aspects necessary to satisfy these requirements.     

复合材料层板低速冲击后剩余压缩强度

对两种材料体系和铺层的复合材料层合板进行低速冲击后压缩强度试验 , 以研究低速冲击后层合板的压缩破坏机理。讨论了表面凹坑深度、 背面基体裂纹长度、 损伤面积以及剩余压缩强度与冲击能量的关系。在试验研究的基础上 , 建立了复合材料低速冲击后剩余强度估算的一种椭圆形弹性核模型。该模型将冲击损伤等效为一刚度折减的椭圆形弹性核 , 采用含任意椭圆核各向异性板杂交应力有限元分析含损伤层合板的应力应变状态 ,并应用点应力判据预测层板的压缩(或压、 剪)剩余强度。理论分析与试验结果对比表明 , 该模型简单有效。      

A coupled approach of optimization,uncertainty analysis and configurational mechanics for a fail‐safe design of structures

In this contribution, a novel method for a fail‐safe optimal design of structures is proposed, which is a coupled approach of optimization employing a genetic algorithm, the structural analysis conducted in the framework of fracture mechanics and uncertainty analysis. The idea of fail‐safe structures is to keep their functionality and integrity even under damage conditions, for example, a local failure of substructures. In the present work, a design concept of a substructure exhibiting a damage accumulating function due to the application of crack arresters is introduced. If such a substructure is integrated within a system of coupled substructures, it will accumulate the damage arising from the boundary conditions change induced by the failure of certain neighbouring structural elements and hinder further damage escalation. The investigation of failure of the damage accumulating substructure is introduced within a finite element framework by a combination of discrete fracturing and configurational mechanics based criteria. In order to design a structure, which will fail safely according to a predefined scenario, uncertainties are taken into account. The developed approach optimizes the configuration of crack arresters within the damage accumulating substructure so that the uncertain crack propagation is hindered and only a local failure of this element occurs. Copyright © 2016 John Wiley & Sons, Ltd.     

Numerical Modeling of the Surface Fatigue Crack Propagation Including the Closure Effect

Presently modeling of surface fatigue crack growth for residual life assessment of structural elements is almost entirely based on application of the Linear Elastic Fracture Mechanics (LEFM). Generally, it is assumed that the crack front does not essentially change its shape, although it is not always confirmed by experiment. Furthermore, LEFM approach cannot be applied when the stress singularity vanishes due to material plasticity, one of the leading factors associated with the material degradation and fracture. Also, evaluation of stress intensity factors meets difficulties associated with changes in the stress state along the crack front circumference. An approach proposed for simulation the evolution of surface cracks based on application of the Strain-life criterion for fatigue failure and of the finite element modeling of damage accumulation. It takes into account the crack closure effect, the nonlinear behavior of damage accumulation and material compliance increasing due to the damage advance. The damage accumulation technique was applied to model the semi-elliptical crack growth from the initial defect in the steel compact specimen. The results of simulation are in good agreement with the published experimental data.