先进复合材料格栅加筋板的总体稳定性分析

通过建立复合材料格栅结构单元的基本力学假设和分析模型, 推导了一种新的等效刚度计算方法, 用于分析复合材料格栅加筋板的总体稳定性。该方法充分考虑了筋条和面板的相互作用和中面偏移效应, 并具有通用性。结合Rayleigh-Ritz能量近似法, 推导出了求解加筋板总体屈曲载荷的通用特征方程; 分析了多种算例, 并与修正平铺法进行了比较, 结果吻合得很好。      

Prediction of post-buckling strength of stiffened laminated composite panels based on the criterion of mixed-mode stress intensity factors

An analytical investigation is conducted to predict the post-buckling strength of laminated composite stiffened panels under compressive loads. When a stiffened composite panel buckles, the skin would deform into a sinusoidal mode shape, and hence induces additional moments and forces near the skin-stiffener interface region. These induced loads would cause the existing small edge delamination cracks to propagate along the skin-stiffener interface, and this in turn would lead to the global failure of the stiffened panel. To reduce the cost of the analytical investigation, the failure of the stiffened panel under post-buckling loads is modeled in two stages: a global analysis to model the post-buckling behavior of the stiffened panel; and a local analysis to model the onset of propagation of the edge delamination crack at the skin-stiffener interface. The results from this study are compared with an experimental investigation conducted by Starnes, Knight, and Rouse (1987). It is found that for the eight different specimens that are considered in this study, the calculated critical energy release rate for the propagation of the edge delamination crack in each specimen differs substantially from those for the others; hence it may be concluded that the total energy release rate would not be a suitable fracture parameter for predicting the post-buckling strength of the stiffened panels. On the other hand, using the fracture criterion based on the critical mixed-mode stress intensity factors, the predicted post-buckling strength of the stiffened panels compares quite favorably with the experimental results and the standard deviation of the error of prediction is less than 10%. Furthermore, by applying the criterion of critical mixed-mode stress intensity factors on a simple damage model, the present analysis is able to predict the significant reduction in the post-buckling strenght of stiffened panels with a damage due to a low-speed impact at the skin-stiffener interface region.This work is supported by ONR, with Dr. Y. Rajapakse as the program official.     

翼肋支撑对复合材料加筋板轴压性能的影响

为确定翼肋支撑对复合材料加筋板轴压性能的影响,对施加翼肋支撑前后的复合材料工型加筋板和帽型加筋板进行压缩试验和数值模拟研究。轴压试验中,通过应变计和影像云纹法实时监测试验件的失稳载荷及失稳模态,通过断面观测分析结构损伤破坏机制。基于ABAQUS软件建立有限元模型模拟加筋板屈曲及后屈曲过程,通过失稳节线及反节线上的应力分布变化分析加筋板破坏机制。计算结果与试验结果相吻合,表明翼肋支撑对不同筋条加筋板失稳模态有影响但均不改变结构失稳载荷,位于节线上的翼肋支撑对工型加筋板破坏载荷影响较小,但位于反节线上的翼肋支撑使帽型加筋板的承载能力提高了26.2%。试验件失稳后应力向反节线上筋条蒙皮界面集中,过高的应力导致界面脱粘,使得结构集中在反节线上破坏。      

加劲钢板剪力墙弹性抗剪屈曲性能研究

系统地分析了三个主要参数(板高厚比λ、肋板刚度比η和柱刚度β)对钢板剪力墙弹性屈曲性能的影响;提出了钢板剪力墙的三种弹性屈曲形式(局部屈曲、整体屈曲和相关屈曲),指出肋板刚度比η=40时(相当于I s=3.7bt3)能保证十字加劲板发生小区格板的局部屈曲或整块板的相关屈曲;指出现行高层建筑钢结构规范用于纵横加劲板局部屈曲荷载的设计表达式仅适用于η≥40的条件,而在η<40范围内,其计算结果偏于不安全;研究了交叉加劲板的弹性屈曲荷载,并与非加劲板和十字加劲板的屈曲性能进行了比较;最终给出了柱刚度对弹性屈曲荷载的修正系数,提出了交叉加劲板弹性屈曲荷载的简化计算公式。     

加强筋截面类型对轨道车辆板件结构隔声影响研究

针对轨道车辆轻量化设计后可能带来的隔声性能降低问题,研究不同截面加强筋铺设对板件隔声性能的改善效果。基于混合有限元-统计能量分析(Hybrid FE-SEA)方法建立轨道车辆加筋板结构隔声特性预测分析模型,系统分析T型、L型、I型和矩形加强筋截面类型对板件隔声性能的影响。研究结果表明,加筋板的刚度和1阶固有频率皆比均质板大,且随加强筋厚度的增大而增大;当加强筋厚度恒定时,T型加筋板的刚度和1阶固有频率最大,L型加筋板次之;敷设厚度15 mm的加强筋,板件的隔声性能最佳;当加强筋的质量、厚度、腹板面积及尺寸、翼板面积相等时,各类型加筋板的计权隔声量Rw差异不大;板件加筋后,刚度控制区的隔声量增幅3 dB~17 dB,1 250 Hz~4 000 Hz中高频段的隔声量增幅1 d B~6 d B。综合分析可知,以计权隔声量为评价标准时,在加强筋质量、腹板面积、翼板面积及尺寸相等时,敷设厚度15 mm加强筋,板件的隔声性能最佳,Rw较均质板可提高1.4 dB~1.5 dB,而加强筋厚度恒定时,T型和L型加筋板的刚度又最佳。相关研究成果可为轨道车辆板件结构加筋优化提供设计参考。     

Optimum buckling design of composite stiffened panels using ant colony algorithm

Optimal design of laminated composite stiffened panels of symmetric and balanced layup with different number of T-shape stiffeners is investigated and presented. The stiffened panels are simply supported and subjected to uniform biaxial compressive load. In the optimization for the maximum buckling load without weight penalty, the panel skin and the stiffened laminate stacking sequence, thickness and the height of the stiffeners are chosen as design variables. The optimization is carried out by applying an ant colony algorithm (ACA) with the ply contiguous constraint taken into account. The finite strip method is employed in the buckling analysis of the stiffened panels. The results shows that the buckling load increases dramatically with the number of stiffeners at first, and then has only a small improvement after the number of stiffeners reaches a certain value. An optimal layup of the skin and stiffener laminate has also been obtained by using the ACA. The methods presented in this paper should be applicable to the design of stiffened composite panels in similar loading conditions.     

Damage tolerant evaluation of cracked stiffened panels under fatigue loading

This paper presents the methodologies for damage tolerant evaluation of stiffened panels under fatigue loading. The two major objectives of damage tolerant evaluation, namely, the remaining life prediction and residual strength evaluation of stiffened panels have been discussed. Concentric and eccentric stiffeners have been considered. Stress intensity factor for a stiffened panel has been computed by using parametric equations of numerically integrated modified virtual crack closure integral technique. Various methodologies for residual strength evaluation, namely, plastic collapse condition, fracture toughness criterion and remaining life approach have been described. Effect of various stiffener sizes and stiffener type (concentric and eccentric stiffeners) on remaining life and residual strength has been studied under constant amplitude load. From the studies, it has been observed that the predicted life is significantly higher with concentric and eccentric stiffener cases compared to the respective unstiffened cases. The percentage increase in life is relatively more in the case of concentric stiffener compared to that of eccentric stiffener case for the same stiffener size and moment of inertia. From the studies, it has also been observed that the predicted residual strength using remaining life approach is lower compared to other methods, namely, plastic collapse condition and fracture toughness criterion and hence remaining life approach will govern the design. It is noted that residual strength increases with the increase of stiffener size.     

Postbuckling analysis of stiffened laminated composite panels,using a higher-order shear deformation theory

The investigation aims at: (i) constructing a modified higher-order shear deformation theory in which Kirchhoff's hypotheses are relaxed, to allow for shear deformations; (ii) validating the present 5-parameter-smeared-laminate theory by comparing the results with exact solutions; and (iii) applying the theory to a specific problem of the postbuckling behavior of a flat stiffened fiber-reinforced laminated composite plate under compression.The first part of this paper is devoted mainly to the derivation of the pertinent displacement field which obviates the need for shear correction factors. The present displacement field compares satisfactorily with the exact solutions for three layered cross-ply laminates. The distinctive feature of the present smeared laminate theory is that the through-the-thickness transverse shear stresses are calculated directly from the constitutive equations without involving any integration of the equilibrium equations.The second part of this paper demonstrates the applicability of the present modified higher-order shear deformation theory to the post-buckling analysis of stiffened laminated panels under compression. to accomplish this, the finite strip method is employed. A C ²-continuity requirement in the displacement field necessitates a modification of the conventional finite strip element technique by introducing higher-order polynomials in the direction normal to that of the stiffener axes. The finite strip formulation is validated by comparing the numerical solutions for buckling problems of the stiffened panels with some typical experimental results.     

复合材料加筋板低速冲击损伤的数值模拟

建立了复合材料加筋板在横向低速冲击载荷作用下的渐进损伤有限元模型。该模型考虑了复合材料加筋板受低速冲击时的纤维断裂、基体开裂及分层脱粘等五种典型的损伤形式, 在层内采用应变描述的失效判据, 结合相应的材料性能退化方案, 通过编写VUMAT用户自定义子程序以实现相应损伤类型的判断和演化。在层间以及筋条与层板间加入界面元, 模拟层间区域的情况, 结合传统的应力失效判据和断裂力学中的能量释放率准则来判断分层损伤的起始和演化规律。通过对数值模拟结果与实验数据的比较, 验证了模型的合理性和有效性。同时探讨了不同位置、不同冲击能量以及含初始损伤(脱粘)等因素对复合材料加筋板低速冲击性能的影响。     

基于临界刚度的复合材料筋材多变量优化设计方法

以船舶结构优化设计为背景,针对目前结构安全余量过高导致加筋板板筋刚度过匹配现状,提出板筋刚度匹配临界刚度的概念,推导了板筋刚度比关系式。以T型复合材料筋材为对象,建立优化模型,基于Isight软件平台对设计变量进行灵敏度分析,简化设计变量。采用多岛遗传算法对筋材开展多变量优化设计,结合工程实际在筋材优化结果基础上确定设计方案,并开展复合材料加筋板力学性能试验研究,验证了多变量优化设计方法的可行性。研究表明:利用提出的加筋板板筋刚度比关系式,可以指导板筋刚度匹配设计;对T型复合材料筋材进行优化设计时,提升腹板高度对优化目标影响最明显;在等刚度约束前提下,提出的T型筋材优化设计方案能够较好地实现优化目标,同时保证了较优的经济性。      

考虑屈曲的复合材料加筋壁板铺层顺序优化

提出了一种考虑屈曲的复合材料加筋壁板铺层顺序优化方法。基于复合材料加筋壁板屈曲载荷求解的能量法,系统推导了轴压载荷作用下复合材料加筋壁板蒙皮、筋条局部屈曲载荷的显示表达式,考虑了加筋壁板各板元之间的弹性支持作用及筋条下缘条的影响,引入工程法求解了加筋壁板整体屈曲载荷。基于国产自主结构分析软件HAJIF中的复合材料铺层工程数据库,以铺层参数为中间变量,利用本文提出的复合材料加筋壁板屈曲载荷求解方法,构建了考虑屈曲的复合材料加筋壁板铺层顺序优化设计流程并完成程序实现,将最小二乘法用于最优铺层顺序与工程铺层数据库的匹配。相比于传统有限元计算方法,本文提出的复合材料加筋壁板屈曲载荷求解方法具备较好的求解精度及求解效率。复合材料加筋壁板优化算例表明,采用本文提出的加筋壁板屈曲载荷分析及其优化方法,在结构重量不变的前提下,屈曲载荷提高约17%,且铺层顺序优化结果可直接从铺层工程数据库中提取并用于工程实际。      

复合材料T型加筋板筋条冲击损伤及冲击后压缩行为试验

本文研究了复合材料加筋板的筋条冲击损伤及冲击损伤对加筋板轴向压缩（CAI）行为的影响。针对T型单筋加筋板,通过落锤法从面板一侧对筋条进行5种能量水平的低速冲击。试验结果表明：冲击筋条产生的面板凹坑不易观察;当冲击能量低于筋条损伤门槛能量时,加筋板筋条无损伤出现,筋条-面板也不会发生脱粘;一旦冲击能量超过筋条损伤门槛能量,筋条的腹板会在弯曲拉伸应力作用下损伤,同时筋条-面板之间会出现严重脱粘。分别对完好和损伤试验件进行压缩试验,试验结果显示：低于门槛能量的冲击对加筋板的压缩屈曲载荷影响不大,同时只会略微降低失效载荷;而冲击造成筋条损伤后,筋条在压缩过程中会由于损伤扩展出现卸载;卸载后的筋条会对面板失去支撑,使面板的屈曲载荷降低,从而大幅地削弱加筋板的承载能力。     

Compressive Behavior of 3D Woven Composite Stiffened Panels: Experimental and Numerical Study

The structural behavior and damage propagation of 3D woven composite stiffened panels with different woven patterns under axial-compression are here investigated. The panel is 2.5D interlock woven composites (2.5DIWC), while the straight-stiffeners are 3D woven orthogonal composites (3DWOC). They are coupled together with the Z-fibers from the stiffener passing straight thought the thickness of the panel. A “T-shape” model, in which the fiber bundle structure and resin matrix are drawn out to simulate the real situation of the connection area, is established to predict elastic constants and strength of the connection region. Based on Hashin failure criterion, a progressive damage model is carried out to simulate the compressive behavior of the stiffened panel. The 3D woven composite stiffened panels are manufactured using RTM process and then tested. A good agreement between experimental results and numerical predicted values for the compressive failure load is obtained. From initial damage to final collapse, the panel and stiffeners will not separate each other in the connection region. The main failure mode of 3D woven composite stiffened panels is compressive failure of fiber near the loading end corner.     

复合材料加筋板的大变形有限元分析

本文将考虑横向剪切变形的Mindlin's理论应用于复合材料加筋板的大变形分析,在Total-Lagrange坐标系下推导了八结点等参弯曲板单元和三结点等参梁单元的增量平衡方程和切线刚度矩阵,非线性问题采用增量法和Newton-Raphson迭代法相结合的方法求解.本文通过一些算例,证明了所采用单元具有良好的收敛性和足够的精度,并讨论了边界条件、纤维铺设角和加筋疏密等因素对复合材加料筋板非线性解的影响.     

多损伤复合材料加筋壁板高周疲劳特性及剩余压缩强度

为研究常用于飞机垂尾的复合材料加筋壁板的冲击疲劳特性,设计了该型加筋壁板多点冲击试验、高周疲劳试验及剩余压缩强度试验。讨论了不同冲击能量对筋条边缘冲击损伤的影响,及施加低应力水平的疲劳载荷后各冲击损伤区域的扩展情况,对比分析了疲劳对冲击后剩余压缩强度的影响。结果表明:40J能量冲击后的损伤面积和凹坑深度较大,C扫描损伤形貌很不规则。100万次低应力疲劳后主损伤区附近衍生出新损伤,导致压缩破坏时产生向上、下夹具扩展的裂痕。该型加筋壁板疲劳后破坏载荷保持率为95.6%,有较好的抗冲击疲劳能力,为加筋壁板耐久性及后屈曲设计提供了思路。     

Impact Analysis of Embedded Delamination Location in Hybrid Curved Laminated Composite Stiffened Panel

Modern, aero structures are predominantly of curved construction characterized by a skin and stiffeners. The latest generation of large passenger aircraft also uses mostly composite material in their primary structure and there is trend towards the utilization of bonding of subcomponents. The presence of delamination is a major problem in composite laminated panels and so, it is of great concern to both the academic and aeronautical industrial worlds Indeed delamination can strongly affect the material strength and, sometimes, can cause their breaking up in service. A Pre-damaged configuration is loaded to study the delamination location and mode for delamination initiation and propagation. A parametric study is conducted to investigate the effect of the location of the delamination propagation when delamination is embedded inbetween plies of the skin-stiffener interface, with the cases i) delamination located at front and inbetween plies of the skin-stiffener interface ii) delamination located in middle and inbetween plies of the skin-stiffener interface iii) delamination located at the end and inbetween plies of the skin- stiffener interface. Further the influence of the location of the delamination on load carrying capacity of the panel is investigated. The effect of location of debonds on crack growth and collapse behavior is analyzed using analysis tool. An analysis tool is applied that includes an approach for predicting interlaminar damage initiation and interlaminar damage growth as well as in-plane damage mechanisms to predict the design of defect free panel.     

Weight optimal design of lateral wing upper covers made of composite materials

The present investigation is devoted to the development of a new optimal design of lateral wing upper covers made of advanced composite materials, with special emphasis on closer conformity of the developed finite element analysis and operational requirements for aircraft wing panels. In the first stage, 24 weight optimization problems based on linear buckling analysis were solved for the laminated composite panels with three types of stiffener, two stiffener pitches and four load levels, taking into account manufacturing, reparability and damage tolerance requirements. In the second stage, a composite panel with the best weight/design performance from the previous study was verified by nonlinear buckling analysis and optimization to investigate the effect of shear and fuel pressure on the performance of stiffened panels, and their behaviour under skin post-buckling. Three rib-bay laminated composite panels with T-, I- and HAT-stiffeners were modelled with ANSYS, NASTRAN and ABAQUS finite element codes to study their buckling behaviour as a function of skin and stiffener lay-ups, stiffener height, stiffener top and root width. Owing to the large dimension of numerical problems to be solved, an optimization methodology was developed employing the method of experimental design and response surface technique. Optimal results obtained in terms of cross-sectional areas were verified successfully using ANSYS and ABAQUS shared-node models and a NASTRAN rigid-linked model, and were used later to estimate the weight of the Advanced Low Cost Aircraft Structures (ALCAS) lateral wing upper cover.     

An analysis methodology for failure in postbuckling skin–stiffener interfaces

Blade-stiffened structures have the potential to produce highly efficient structures, particularly when the large strength reserves available after structural buckling, in the postbuckling range, are exploited. In experimental tests of postbuckling stiffened structures made from fibre-reinforced composites, failure typically initiates at the interface of the skin and stiffener and leads to rapid and even explosive failure. A methodology has been developed for analysing collapse in postbuckling composite structures that involves predicting the initiation of interlaminar damage in the skin–stiffener interface. A strength-based criterion is monitored in each ply using a local model of the skin–stiffener interface cross-section. For the analysis of large structures, a global analysis is first run to obtain the complete postbuckling deformation field, which is then input onto a local model using a global–local analysis technique. The coordinates of the local model can easily be moved to rapidly assess failure initiation at numerous skin–stiffener interface locations throughout the global structure. The analysis methodology is compared to experimental results for two-dimensional T-section specimens and large, fuselage-representative stiffened panels and is shown to give accurate predictions of the failure load and failure mechanisms. The use of the approach for the analysis of postbuckling composite structures has application for the design and certification of the next generation of aircraft.     

Residual strength analyses of stiffened and unstiffened panels––Part II: wide panels

This paper highlights the results from fracture analyses conducted on the FAA/NASA wide panels (with and without stiffeners) using structural analysis of general shells code and the critical crack-tip-opening angle (CTOA) fracture criterion. The critical CTOA and plane-strain core height values, calibrated from a fit to the experimental load-against-crack-extension results from a series of unstiffened panels (76–1016 mm wide) tests with anti-buckling guides (Part I of this paper), were used in the analyses of wide stiffened and unstiffened panels. As discussed in Part I of this paper, high constraint around the crack front like plane strain has been accounted for by using the “plane-strain core” option in all analyses. By accounting for high constraint around crack front, it was possible for the critical CTOA fracture criterion to predict wide panel failures from small laboratory tests. As followed in Part I of this paper, rivet flexibility and stiffener failures in the analyses of wide panels were based on methods and criteria like that currently used in the industry. Analyses were able to predict stable crack growth and residual strength of both stiffened and unstiffened panels with various amounts of multiple-site damage within ±10% of the test results. Finally, it has been demonstrated that, it is possible to predict the residual strength of wide stiffened and unstiffened panels with critical CTOA calibrated from small laboratory coupons.     

Analysis of Mode I and Mode II Crack Growth Arrest Mechanism with Z-Fibre Pins in Composite Laminated Joint

This paper presents the numerical study of the mode I and mode II interlaminar crack growth arrest in hybrid laminated curved composite stiffened joint with Z-fibre reinforcement. A FE model of hybrid laminated skin-stiffener joint reinforced with Z-pins is developed to investigate the effect of Z- fibre pins on mode I and mode II crack growth where the delamination is embedded inbetween the skin and stiffener interface. A finite element model was developed using S4R element of a 4-node doubly curved thick shell elements to model the composite laminates and non linear interface elements to simulate the reinforcements. The numerical analyses revealed that Z-fibre pinning were effective in suppressing the delamination growth when propagated due to applied loads. Therefore, the Z-fibre technique effectively improves the crack growth resistance and hence arrests or delays crack growth extension.