Structural fatigue and joint degradation

This paper deals with three subjects on fatigue behavior of single lap joint structures based on the test results to discuss the structural integrity problems of aging aircraft. Firstly, the effect of the following factors on fatigue life were examined; fastener type, fastener row, squeezing force, overload, underload as well as corrosion. Secondly, the simulation problems of fatigue behavior of fuselage structures by uni-axial laboratory test using stiffened panel specimen are discussed. Thirdly, the effects of service load components on fatigue behavior of fuselage structure are examined using 1/3 scale model of a B-737 aircraft. Model specimens were tested by pneumatic cycles with and without synchronized bending. Test results suggested that fatigue behavior of fuselage structure can be estimated by evaluating the largest principal stress under complex stress conditions.     

Integration of Mechanics and Acoustics in a Sandwich Fuselage. Part IV

Until now only the stiffened skin structural concept has been discussed. A different structural concept is the sandwich concept. Sandwiches consist out of layers. The outer layers are called facings and are generally thin and of high density. These facings are supposed to resist most of the edgewise loads and flat-wise bending moments. The inner layer is called the core and is generally rather thick and of low density. The task of the core is to separate and stabilize the two facings, transmit shear between the facings and provide most of the shear rigidity. For sandwich panels no stiffeners are needed. Therefore no mass will be lost in stiffeners resulting in a relative high value of mass per unit area of the skin which results in a better TL according to the mass law. Also the core can be made of a material with high insulation properties (acoustic and thermal). The number of discrete stiffeners can then be minimized, since they are only required at places where high concentrated forces have to be introduced (wing, landing gear, etc.) or diverted (from cut-outs). This can reduce the production and maintenance cost. So it can be concluded that the sandwich concept offers great potential for multidisciplinary fuselage design. In this part the integration of structural and acoustical aspects will be discussed. First the structural aspect will be discussed followed by the acoustical aspect. Finally the possibilities to integrate these aspects are explained.     

Some computational tools for the analysis of through cracks in stiffened fuselage shells

Reliable analytical methods that predict the structural integrity and residual strength of aircraft fuselage structures containing cracks are needed to help to understand the behavior of pressurized stiffened shells with damage, so that it becomes possible to determine the safe life of such a shell. Of special importance is the ability to determine under what conditions local failure, once initiated, will propagate. In this paper we shall present a reliable and efficient method for computing the energy release rate for cracks of varying length in a typical stiffened metallic fuselage under general loading conditions. The models used in the simulation are derived from an extensive analysis of a fuselage barrel section subjected to operational flight loads. Energy release rates are computed as a function of the length of the crack, its location, and the crack propagation mode.     

爆炸荷载作用下复合材料加筋板的动力响应

为了减轻抗爆结构质量,采用玻璃纤维增强聚合物基复合材料（SMC）与碳纤维增强聚合物基复合材料（CFRP）预浸料,通过数值模拟和等效计算理论,对传统加筋抗爆板结构进行轻质高强设计。利用LS-DYNA有限元数值模拟软件进行分析,发现在爆炸荷载作用下加筋板的运动以弹性运动为主,该种复合材料具有较好的抗爆性能。对复合材料加筋板结构进行参数化分析,发现在爆炸荷载作用下横筋对加筋板结构最大位移值影响最大,纵筋和面板对加筋板的影响依次减小。结合刚度折算方法,建立了爆炸荷载作用下正交异性加筋板结构动力响应分析理论。利用该理论计算得到板结构在爆炸荷载作用下的最大位移,与数值模拟对比发现两者结果较为接近,为加筋抗爆板的设计提供了一种简化有效的计算方法。     

Impact Force Identification of CFRP Structures Using Experimental Transfer Matrices

This paper presents a method for identifying the location and force history of an impact force acting on CFRP structures such as laminated plates and stiffened panels. The identification method is an experimental one without using any analytical model of the structure. Here, experimental transfer matrices, which relate the impact force to the corresponding responses of PZT sensors, are used to identify the impact force. The transfer matrices are preliminarily constructed from the measured data obtained by impact tests with an impulse hammer. To identify the impact location, the arrival times of the flexural waves to the PZT sensors are used, and an analog band-pass filter is used to obtain waves with a specified frequency. The wave velocity is determined experimentally from impact test results. The present method is verified experimentally by performing impact force identification of CFRP laminated plates and CFRP stiffened panels. The results reveal that the location and force history of the impact force can be identified accurately and rapidly using the present method.     

Sound Transmission Loss Prediction of the Composite Fuselage with Different Methods

Increase of sound transmission loss(TL) of the fuselage is vital to build a comfortable cabin environment. In this paper, to find a convenient and accurate means for predicting the fuselage TL, the fuselage is modeled as a composite cylinder, and its TL is predicted with the analytical, the statistic energy analysis (SEA) and the hybrid FE&SEA method. The TL results predicted by the three methods are compared to each other and they show good agreement, but in terms of model building the SEA method is the most convenient one. Therefore, the parameters including the layup, the materials, the geometry, and the structure type are studied with the SEA method. It is observed that asymmetric laminates provide better sound insulation in general. It is further found that glass fiber laminates result in the best sound insulation as compared with graphite and aramid fiber laminates. In addition, the cylinder length has little influence on the sound insulation, while an increase of the radius considerably reduces the TL at low frequencies. Finally, by a comparison among an unstiffened laminate, a sandwich panel and a stiffened panel, the sandwich panel presents the largest TL at high frequencies and the stiffened panel demonstrates the poorest sound insulation at all frequencies.     

Post-buckling behaviour of flat stiffened composite panels: Experiments vs. analysis

The paper is focused on the development of a validated procedure for modelling, by means of Finite Element tools, the post-buckling behaviour of stiffened composite flat panels subjected to compression loads. The experimental data for model validation were collected during a test campaign on two sets of CFRP flat stiffened panels.     

Two‐stage fatigue life evaluation of an aircraft fuselage panel with a bulging circumferential crack and a broken stringer

The article presents two‐stage fatigue life evaluation of a stiffened aluminium aircraft fuselage panel, subject to ground–air–ground pressure cycles, with a bulging circumferential crack and a broken stringer. As a worst‐case scenario, it is assumed that double cracks start at the edge of a rivet hole both in the skin and in the stringer simultaneously. In the first stage, fatigue crack growth analysis is performed until the stringer is completely broken with the crack on the fuselage skin propagating. After the stringer is completely broken, the effect of bulging crack on the fatigue life of the panel is investigated utilizing the stress intensity factors determined by the three‐dimensional finite element analyses of the fuselage panel with the broken stringer. It is concluded that bulging of the skin due to the internal pressure can have significant effect on the stress intensity factor, resulting in fast crack propagation after the stringer is completely broken.     

Erichsen cupping test on thermosetting CFRP sheets

The deformation of carbon-fiber-reinforced plastic (CFRP) sheets consisting of a thermosetting resin and continuous fibers has been investigated. The bending of CFRP sheets was achieved under a suitable forming temperature and strain path. Formability indexes and a forming limit diagram (FLD) are indispensable data for showing the process window of CFRP sheets. There are four formability indexes: bendability, deep drawability, stretchability, and stretch-flangeability. Bendability and deep drawability can be evaluated by conducting a stretch-bending test, and stretchability can be evaluated by conducting an Erichsen cupping test. If the process window can be predicted, the range of applications using CFRP can be expanded. In this study, the Erichsen index of CFRP sheets indicates the stretchability of laminated CFRP structures.     

Repair of Composites: Design Choices Leading to Lower Life-Cycle Cost

The fabrication cost of composite aircraft structures is revisited and the effect of part size on cost is examined with emphasis on design decisions which affect the ease of (bonded) repair and the total cost of the part and subsequent repairs. The case of moderately loaded stiffened fuselage or wing panels under compression is analysed in detail and the fabrication cost of the panel made as a single piece or as an assembly of smaller identical components or modules is determined. The cost of special purpose repairs for two different damage sizes is compared to removing and replacing damaged modules. Hand layup and automated processing are compared. It is found that for certain repair sizes removing and replacing modules leads to lower overall cost as compared to applying a special purpose repair.     

钢材与复合材料多形式加筋平板透声性能消声水池实测分析

以复合材料声呐导流罩加筋板结构多目标综合优化为工程背景,开展消声水池模型插入损失测试,横向对比分析钢制桁架筋材与四种复合材料筋材加筋板近场透声特性,综合讨论了筋材材质、筋材结构形式以及入射角度对加筋板近场透声性能的影响。测试及分析结果表明:桁架式加筋板近场透声性能优异且稳定;细矮型筋材可提高加筋板整体透声性能。研究结果为声呐导流罩复合材料加筋板结构的优化设计提供了参考。     

Blast response of cracked steel box structures repaired with carbon fibre-reinforced polymer composite patch

In this paper the blast resistance of cracked steel structures repaired with fibre-reinforced polymer (FRP) composite patch are investigated. The switch box which has been subjected to blast loading is chosen to study. The steel material is modelled using isotropic hardening model, pertaining to Von Mises yield condition with isotropic strain hardening, and strain rate-dependent dynamic yield stress based on Cowper and Symonds model. Three different cracked structures are chosen to investigate their capability in dissipating the blast loading. To improve the blast resistance, the cracked steel structures are stiffened using carbon fibre-reinforced polymer (CFRP) composite patches. The repaired patches reduce the stress field around the crack as the stress is transferred from the cracked zone to them. This situation prevents the crack from growing and extends the service life of the steel structure. It will be shown that CFRP repairing can significantly increase the blast resistance of cracked steel structures.     

Impact behaviour of stiffened CFRP sections

Dropweight impact tests have been performed on thin CFRP panels stiffened with blade or T-stiffeners and comparisons made with similar plain panels. The change in structural response of the panels is governed by the amount of damage sustained during impact. The increase in panel stiffness is associated with the suppression of backface cracking but larger areas of delamination.     

On the analysis of composite structures with material and geometric non-linearities

If major weight saving is to be realised it is essential that composites be used in “primary” structural components, i.e., wing and fuselage skins. To this end it is essential that analytical tools be developed to ensure that composite structures meet the FAA damage tolerance certification requirements. For stiffened composite panels one potential failure mechanism is the separation of the skin from the stiffeners; resulting from excessive “through the thickness” stresses. This failure mechanism is also present in bonded composite joints and composite repairs. Currently failure prediction due to in-plane loading appears to be relatively well handled. Unfortunately, this is not yet true for matrix-dominated failures. Consequently, it is essential that a valid analysis methodology capable of addressing all of the possible failure mechanisms, including failure due to interlaminar failure, be developed. To aid in achieving this objective the present paper outlines the results of a series of experimental, analytical and numerical studies into the matrix-dominated failures of rib stiffened structures.     

Detailed design of a lattice composite fuselage structure by a mixed optimization method

In this article, a procedure for designing a lattice fuselage barrel is developed. It comprises three stages: first, topology optimization of an aircraft fuselage barrel is performed with respect to weight and structural performance to obtain the conceptual design. The interpretation of the optimal result is given to demonstrate the development of this new lattice airframe concept for the fuselage barrel. Subsequently, parametric optimization of the lattice aircraft fuselage barrel is carried out using genetic algorithms on metamodels generated with genetic programming from a 101-point optimal Latin hypercube design of experiments. The optimal design is achieved in terms of weight savings subject to stability, global stiffness and strain requirements, and then verified by the fine mesh finite element simulation of the lattice fuselage barrel. Finally, a practical design of the composite skin complying with the aircraft industry lay-up rules is presented. It is concluded that the mixed optimization method, combining topology optimization with the global metamodel-based approach, allows the problem to be solved with sufficient accuracy and provides the designers with a wealth of information on the structural behaviour of the novel anisogrid composite fuselage design.     

Manufacture of Composite Panels with J-Shape Stiffeners

The use of composite materials in primary structure of aircraft is becoming more common. Stiffened composite panels have been designed for a variety of load carrying conditions in aircraft. In this, study, manufacturing techniques for stiffened composite panels were reviewed and discussed. Co-curing and secondary bonding manufacturing techniques were used to produce four stiffened composite panels. Experience gained with these techniques is discussed. Ultrasonic C-scan inspection was carried out to detect any defects in the panels. These panels were tested in compression to verify their structural capability in the post-buckled state. This paper presents the details of the tool design concepts and manufacturing techniques, and identifies and discusses the advantages and disadvantages of each manufacturing technique.     

Cyclic fatigue delamination of carbon fiber-reinforced polymer-matrix composites: Data analysis and design considerations

Activities toward standardization of fracture mechanics tests on carbon fiber-reinforced polymer-matrix (CFRP) composites have recently focused on cyclic fatigue under mode I (tensile opening), mode II (in-plane shear) and mixed-mode I/II loading. Data from recent round robins performed by Technical Committee 4 (TC4) of the European Structural Integrity Society (ESIS) and from preliminary testing of additional CFRP epoxy laminates at the authors’ laboratories are analyzed with different approaches in attempts to reduce scatter and to identify parameters for CFRP structural design. Selected test data comparing load and displacement control for the cyclic fatigue tests are also discussed. Specifically, threshold values from Paris-law data fitting are compared with values from fitting with a modified Hartman–Schijve approach. Independent of the approach used for the analysis, mode I threshold values of selected CFRP seem to be in the range between about 30 and 100 J/m², i.e., roughly around the range of critical mode I energy release rate values (denoted by G_IC) obtained from fracture testing of neat commercial epoxy resins, but clearly below quasi-static initiation G_IC-values for unidirectional CFRP composites. Implications for CFRP structural design based on mode I fatigue fracture mechanics test data are briefly discussed.     

T型截面多级加筋柱壳的缺陷敏感性及优化研究

该文基于非线性显式动力学方法进行后屈曲分析,获得了四种加筋柱壳(均匀加筋柱壳、矩形截面双向多级加筋柱壳、T型截面单向和双向多级加筋柱壳)整个后屈曲过程的轴压位移-载荷曲线,并以模态缺陷为例,比较了四者的缺陷敏感性,结果显示T型截面双向多级加筋柱壳呈现出显著的低缺陷敏感性和较强的可设计性。该文还将缺陷敏感性分析结果与对应的完善结构后压溃稳定平衡路径进行了对比,发现两者一定程度上的趋势一致性。这表明对完善结构运用显式动力学方法进行单次后屈曲分析,即可同时获得其承载能力和缺陷敏感性,大大减少计及缺陷敏感性结构设计的计算量。最后,该文进一步构造了面向低缺陷敏感性的T型截面多级加筋柱壳优化模型,算例表明该方法可以高效地获得可靠性更强的优化解。     

Application of the cohesive model for predicting the residual strength of a large scale fuselage structure with a two-bay crack

The residual strength of a curved and stiffened panel containing a two-bay crack was assessed using the cohesive model. This panel represents a section of a wide-body aeroplane fuselage. The tests were conducted at IMA GmbH Dresden in cooperation with Airbus Industries Germany. The structural panel was modelled using 3D finite elements and a layer of cohesive elements ahead of each crack tip allowing for 70 mm crack extension. Identification of the cohesive parameters was done on small laboratory test pieces. Special effort was made for the transfer of these parameters to the structure. Reasonably conservative predictions of the residual strength of the panel were achieved. The boundary conditions of the loading devices of the test rig are shown to have substantial influence on the predictions.     

On damage tolerance design of fuselage structure (longitudinal cracks)

Aircraft structure is the most obvious example where functional requirements demand light weight and, therefore, high operating stresses. An efficient structural component must have three primary attributes; namely, the ability to perform its intended function, adequate service life and the capability of being produced at reasonable cost. To ensure the safety of aircraft structures, the Air Force requires damage tolerance analysis. This paper focuses its attention on designing a fail-safe fuselage structure. Two types of damage most frequently associated with the structural integrity of the fuselage are longitudinal cracks under high hoop stresses induced by cabin pressurization and circumferential cracks under stresses from vertical bending of the fuselage. The analysis of these types of cracks is complex, first due to the complex structural configuration (i.e. frames, skin longeron and crack stopper straps) and secondly due to the influence of the curvature of the shell. Various analytical and empirical approaches have been used to study the damage tolerance capability of the fuselage structure. Due to the lack of a comprehensive model to calculate the stress intensity factors for the complex structure, experiments usually have been performed to measure the crack growth rates and to demonstrate the residual strength of fuselage-type structural components containing circumferential and longitudinal cracks.

In this paper various analytical and empirical approaches used in evaluating the damage tolerance capability of the fuselage structure are critically evaluated and compared. A model which accounts for the influence of frames, straps and curvature is developed. This model is then used in an example problem having typical military cargo aircraft fuselage structural elements. The Air Force damage tolerance requirements are discussed in detail.