钻孔分层损伤对复合材料层合孔板压缩力学行为的影响

钻孔分层损伤对复合材料层合孔板的承载能力和失效模式有着显著的影响。通过实验和仿真相结合的方式,开展单一预制分层缺陷下、双分层缺陷同侧耦合及双分层缺陷异侧耦合作用下复合材料层合孔板的压缩承载能力及失效模式的研究。通过预埋聚四氟乙烯薄膜,制备了含单一圆形预制分层缺陷的碳纤维增强树脂复合材料开孔板试件,采用浸没式超声C扫和数字图像DIC技术分别对复合材料层合板损伤和法向变形进行检测,研究含不同尺寸预制分层开孔层合板在压缩载荷下的分层扩展及失效变形特征,进而揭示分层缺陷大小对其承载能力的影响机制。构建基于内聚力单元方法的含孔复合材料层合板数值模型,对比实验修正模型,探索了单一预制分层缺陷下碳纤维增强树脂复合材料开孔板的损伤扩展机制,并在此模型基础上开展双分层缺陷耦合作用下复合材料开孔板在压缩载荷作用下的屈曲变形、分层扩展和承载能力的数值预测和分析。实验结果表明：含单一圆形预制分层缺陷的碳纤维增强环氧树脂复合材料开孔层合板试件呈现出初始受压、局部屈曲、整体屈曲后破坏的失效模式,预制分层缺陷对复合材料孔板压缩力学性能有显著影响,随着缺陷的增大压缩承载能力逐渐下降。双分层缺陷耦合作用数值分析表明：双...     

复合材料层合板剪切稳定性试验及强度预测

对无损伤及含冲击损伤的复合材料层合板进行了剪切稳定性试验,基于数字图像相关方法 (Digital image correlation,DIC)对层合板屈曲后屈曲行为进行了实时测量。试验结果表明：引入冲击损伤后,复合材料层合板剪切屈曲波形、屈曲载荷无明显变化,失效模式转变,承载能力下降了9.69%。随后,基于断裂面失效理论,建立了考虑剪切非线性效应的复合材料渐进损伤失效模型,并对复合材料层合板剪切失效过程进行了模拟。模型采用软化夹杂法将冲击损伤等效简化,直接将损伤区的几何边界信息写入材料模型中,不需要对冲击损伤区进行切割,从而保证了整体网格质量。与试验结果对比发现：模型考虑剪切非线性对屈曲载荷预测无明显影响,对后屈曲承载能力的预测精度影响较大,不考虑剪切非线性效应时的误差可达20%以上;软化夹杂法可以有效地模拟冲击损伤,预测的含冲击损伤的复合材料层合板的屈曲载荷、破坏载荷误差分别为-3.17%、-1.27%。     

外突褶皱及其打磨处理对层合板压缩失效行为影响

评估打磨处理对含外突褶皱结构承载能力的影响，是开展复合材料制造质量评定与制定装配策略的重要依据。本文采用实验与数值模拟相结合的方法研究了外突褶皱及打磨处理对层合板压缩失效行为的影响。实验方面，借助数字图像DIC技术监测加载过程表面应变分布，采用电子显微相机捕捉损伤过程，分析试样的压缩承载行为和破坏模式。数值模拟方面，采用Hashin失效准则与内聚力方法构建高保真度含褶皱层合板层内/层间失效分析有限元模型，探讨含外突褶皱及打磨褶皱层合板的应力分布特征及失效机制并分析褶皱高度对层合板压缩强度影响。结果表明：外突褶皱降低层合板压缩承载能力，而打磨处理会进一步降低层合板的承载能力与刚度。本文所建立的数值模型与实验结果吻合较好。对于外突褶皱层合板，层合板向褶皱凸起一侧屈曲，褶皱上方发生分层损伤并向端部扩展，纤维损伤从褶皱下侧铺层波谷处向其他层扩展；对于打磨褶皱层合板，主层板向褶皱凸起的相反方向屈曲，首先在打磨断层处发生面外拱起分层损伤，随后褶皱上方纤维层发生拱起断裂，纤维损伤向其他层扩展。随外突褶皱高度的增加，含褶皱层合板及打磨褶皱层合板的压缩失效载荷均显著降低。以上研究可为含外突褶皱复合材料结...     

含纤维波纹缺陷复合材料层合板的损伤分析

纤维波纹是复合材料层合板制备过程中的一种常见缺陷,会导致其刚度和强度显著下降,有效地预测含波纹缺陷复合材料层合板的失效强度具有显著的意义。基于此,本文采用解析的方式分别构造了纤维波纹呈正弦起伏与余弦起伏状的复合材料层合板模型。利用该模型,以Tsai-Wu准则作为失效判据,研究了一种含纤维波纹的碳纤维/环氧树脂复合材料层合板在受压情况下的损伤演化过程,得到了碳纤维/环氧树脂复合材料层合板的初始损伤强度。与有限元方法计算得到的损伤位置和损伤强度非常吻合,验证了本文算法的正确性。另外,相比于有限元方法,本文所述计算方法具有模型构造简单、计算效率高等优点,便于快速分析和确定含纤维波纹缺陷复合材料层合板的损伤位置与损伤强度。      

基于超弹性特性的SMA复合材料层合板低速冲击损伤数值模拟方法

基于ABAQUS有限元软件结合VC++6.0程序设计,建立了含不同铺层角度、不同排列密度形状记忆合金(SMA)纤维的复合材料层合板有限元模型。将基于Brinson本构模型的SMA分段线性超弹性模型以及判断复合材料层内失效的三维HASHIN失效准则编译至ABAQUS/VUMAT子程序,使用界面单元模拟复合材料层间区域,建立了SMA复合材料层合板的低速冲击损伤及冲击后剩余强度数值模拟方法。对比了不含SMA纤维层合板、含SMA纤维层合板、含普通金属丝层合板在不同冲击能量下的损伤响应。进一步分析了SMA纤维体积分数和直径变化对冲击响应的影响。冲击后剩余压缩强度模拟结果表明:冲击能量为16J时,含体积分数25%、直径0.5mm的SMA纤维层合板的冲击后剩余压缩强度相比不含SMA纤维层合板提高5.78%、相比含普通金属丝层合板提高4.69%。随着SMA纤维体积分数提高,层合板的抗低速冲击能力增强,当体积分数一定时,较细的(0.3mm)SMA纤维比粗的(0.6mm)SMA纤维对层合板的抗低速冲击能力增强效果更好。     

低速冲击后复合材料层合板的压缩破坏行为

对缝纫层合板和无缝纫层合板进行低速冲击后压缩破坏实验,以研究低速冲击后层合板的压缩破坏机理。采用C扫描、X射线、热揭层等技术对层合板内的损伤进行测量和对比。结果表明,界面不是很强的碳纤维增强复合材料层合板低速冲击后受压时,层合板非冲击面的子层屈曲及其扩展是导致层合板冲击后压缩强度下降的重要因素,而且子层屈曲主要是沿垂直载荷的方向(90°)扩展;对于准各向同性板,屈曲子层中与母层相邻的铺层的方向一般为90°。层合板的剩余压缩强度与板的冲击损伤面积无直接关系。      

低速冲击作用下碳纤维复合材料铺层板的损伤分析

建立了一个有效计算模型, 以分析碳纤维复合材料层合板在低速冲击作用下的层内和层间失效行为。针对铺层板的层内损伤, 在基于应变描述的Hashin 失效准则的基础上, 建立了单层板的逐渐累积损伤分析模型;针对铺层板的脱层损伤, 建立了各向同性脱层损伤模型, 通过结合传统的应力失效准则和断裂力学中的能量释放率准则定义了界面损伤演化规律, 并在潜在产生脱层的区域模拟为粘结接触, 并将脱层损伤模型作为界面的接触行为。该计算模型通过商用有限元软件ABAQUS/ Explicit 的用户子程序实现。使用该计算模型对碳纤维增强环氧树脂复合材料层合板在横向低速冲击作用下的损伤和变形行为进行预测分析。数值仿真的结果与试验结果进行了比较, 取得了满意的结果, 验证了该模型的正确性。      

基于微观力学失效理论的CFRP多向层合板低速冲击行为预测

基于微观力学失效(MMF)理论对碳纤维增强复合材料(CFRP)多向层合板在低速冲击载荷下失效机制及损伤过程进行分析和预测。建立基于MMF理论的层合板结构冲击损伤行为分析方法。首先, 使用MMF理论对冲击过程中组分的失效类别进行判别; 然后, 根据组分失效的类别制定出相应的材料性能退化方案来实现对复合材料在低速冲击下的逐步失效分析;在ABAQUS平台上开发了基于显示分析的用户材料子程序(VUMAT), 即基于MMF理论的层合板冲击损伤分析程序;最后, 利用MMF理论冲击损伤行为分析方法, 对UTS50/E51碳纤维增强复合材料多向层合板在小能量低速冲击情况下的失效机制和损伤形貌进行预测, 并将预测结果与试验结果进行对比, 分析了利用MMF理论预测冲击损伤这一方法的准确性。结果表明理论预测的凹坑直径与试验测试的凹坑直径误差为4.8%, 预测的失效机制和损伤形貌与实际观察的一致。      

剪切载荷作用下复合材料挖补修理层合板试验及有限元分析

为了确定剪切载荷作用下含非穿透损伤复合材料挖补修理层合板的破坏模式和抗剪切能力,进行了复合材料挖补修理层合板的剪切试验,并与未损伤复合材料层合板进行对比。试验结果表明,复合材料挖补修理后的层合板具有较高的强度恢复率,且不影响层合板的后屈曲承载能力。同时,建立了剪切载荷作用下复合材料挖补修理层合板的有限元分析（FEA）模型,复合材料母板和补片采用了三维Hashin准则来判定材料失效,母板层与层之间采用零厚度界面单元以有效模拟剪切载荷作用下复合材料母板上、下子板之间的分层。该模型得到的破坏模式与试验结果基本相符。由于挖补修理的设计与工艺复杂性,理论模拟的破坏载荷与试验结果虽不能完全吻合,但其最大15%左右的差异能够满足修理设计的需要。以上结果说明,该模型对剪切载荷作用下复合材料挖补修理层合板的破坏模式和破坏载荷能够进行工程适用的预测。     

纤维/镁合金混杂层合板低速冲击响应及损伤模拟

为了研究新型纤维增强镁合金混杂层合板在低速冲击下的力学响应,分别对由玻璃纤维、碳纤维和二者混杂增强的AZ31B镁合金层合板在不同冲击能量下的落锤低速冲击试验进行了数值模拟。基于镁合金各向异性塑性本构和指数关系界面脱粘内聚力本构模型,同时纤维复合材料层采用三维Hashin失效准则且引入刚度折减,编写了复合材料层板损伤的VUMAT子程序,并将该子程序嵌入ABAQUS/Explicit中实现对层合板冲击过程的模拟。研究了该纤维层合板在不同冲击能量下的动态冲击响应以及脱粘与损伤演化规律,分析了冲击载荷、形变和能量吸收随时间的变化规律。模拟结果表明:在冲击能较小时,首先在冲击背面出现基体开裂,随着冲击能的增加,层合板受冲击面出现由无明显损伤到出现基体开裂和纤维断裂的现象;与单一碳纤维增强的镁合金层合板复合材料相比,单一玻璃纤维增强的镁合金层合板在冲击载荷作用时能够吸收更多的能量,碳纤维层内混杂合适的玻璃纤维铺层能够提高碳纤维增强镁合金层合板的抗冲击性能。     

Damage growth analysis of low velocity impacted composite panels

Low velocity impact loading in aircraft composite panels is a matter of concern in modern aircraft and can be caused either by maintenance accidents with tools or by in-flight impacts with debris. The consequences of impact loading in composite panels are matrix cracking, inter laminar failure and, eventually, fiber breakage for higher impact energies. Even when no visible impact damage is observed on the surface at the point of impact, matrix cracking and inter laminar failure can occur, and the carrying load of the composite laminates is considerably reduced. The greatest reduction in loading is observed in compression due to laminae buckling in the delaminated areas.

The objective of this study is to determine the limit loading capacity and the damage growth mechanisms of impacted composite laminates when subjected to compression after impact loading. For this purpose a series of impact and compression after impact tests were carried out on composite laminates made of carbon fiber reinforced epoxy resin matrix. Four stacking sequences representative of four different elastic behaviours were used. Results show that the compressive, after impact, failure stress is influenced by the stacking sequence but a relatively independent strain to failure is observed.     

Failure mechanisms on composite specimens subjected to compression after impact

Composite panels are widely used in aeronautic and aerospace structures due to their high strength/weight ratio. The stiffness and the strength in the thickness direction of laminated composite panels is poor since no fibres are present in that direction and out-of-plane impact loading is considered potentially dangerous, mainly because the damage may be left undetected. Impact loading in composite panels leads to damage with matrix cracking, inter-laminar failure and eventually fibre breakage for higher impact energies. Even when no visible impact damage is observed at the surface on the point of impact, matrix cracking and inter-laminar failure can occur, and the carrying load of the composite laminates is considerably reduced. The greatest reduction in loading is observed in compression due to laminae buckling in the delaminated areas. The objective of this study is to determine the mechanisms of the damage growth of impacted composite laminates when subjected to compression after impact loading. For this purpose a series of impact and compression after impact tests were carried out on composite laminates made of carbon fibre reinforced epoxy resin matrix. An instrumented drop-weight-testing machine and modified compression after impact testing equipment were used together with a C-scan ultrasonic device for the damage identification. Four stacking sequences of two different epoxy resins in carbon fibres representative of four different elastic behaviours and with a different number of interfaces were used. Results showed that the delaminated area due to impact loading depends on the number of interfaces between plies. Two buckling failure mechanisms were identified during compression after impact, which are influenced more by the delamination area than by the stacking sequence.     

Damage Tolerance of Pre-Stressed Composite Panels Under Impact Loads

An experimental test campaign studied the structural integrity of carbon fibre/epoxy panels preloaded in tension or compression then subjected to gas gun impact tests causing significant damage. The test programme used representative composite aircraft fuselage panels composed of aerospace carbon fibre toughened epoxy prepreg laminates. Preload levels in tension were representative of design limit loads for fuselage panels of this size, and maximum compression preloads were in the post-buckle region. Two main impact scenarios were considered: notch damage from a 12 mm steel cube projectile, at velocities in the range 93–136 m/s; blunt impact damage from 25 mm diameter glass balls, at velocities 64–86 m/s. The combined influence of preload and impact damage on panel residual strengths was measured and results analysed in the context of damage tolerance requirements for composite aircraft panels. The tests showed structural integrity well above design limit loads for composite panels preloaded in tension and compression with visible notch impact damage from hard body impact tests. However, blunt impact tests on buckled compression loaded panels caused large delamination damage regions which lowered plate bending stiffness and reduced significantly compression strengths in buckling.     

A study on low velocity impact damage evaluation and repair technique of small aircraft composite structure

Composite structures are very prone to damage at fairly modest levels of impact energy due to foreign object damages. A repair technique using external patch is recognized as an effective method to recover the damaged structures during service life. This work is focusing on the impact damage evaluation and the external patch repair techniques of the aircraft composite structure. The impact damages of composite laminates of the carbon/epoxy UD laminate and the carbon/epoxy fabric face sheets-honeycomb core sandwich laminate are simulated by the drop-weight type impact test equipment. The damaged specimens are repaired using the external patch repair method after removing the damaged area. The compressive strength test and analysis results of the repaired impact damaged specimens are compared with the compressive strength test and analysis results of the undamaged specimens and the impact damaged specimens. Finally, the strength recovery capability after repairing is investigated.     

低速冲击下复合材料层板压缩许用值

为评估航空结构中常用的T300级和T800级2种碳纤维/环氧树脂复合材料层压板的冲击后压缩许用值,对2种材料体系下具有不同厚度及铺层的层板进行了低速冲击和冲击后压缩试验;讨论了冲击能量、凹坑深度、损伤面积及冲击后剩余压缩强度等之间的关系,以及厚度、铺层、表面防护等因素对其造成的影响;重点关注了2种材料体系下各组层板的目视勉强可见冲击损伤(BVID)形成条件以及含BVID层板的剩余强度.结果表明:厚度及铺层对层板的凹坑深度-冲击能量关系影响较大,而对冲击后压缩强度-凹坑深度及冲击后压缩破坏应变-凹坑深度关系影响较小,且在相同铺层比例下,BVID对应的冲击能量随厚度近似呈线性增长.X850层板的损伤阻抗性能明显优于CCF300/5228层板的,但二者损伤容限性能相当.加铜网、涂漆等表面处理显著提高了层板的损伤阻抗,但对损伤容限性能影响不大;在损伤不超过BVID时,所有CCF300/5228试件的压缩破坏应变均大于4 000 με,而X850材料体系下压缩破坏应变均在3 000 με之上.     

Experimental determination of elastic properties of impact damage in carbon fibre/epoxy laminates

This paper describes an investigation of in-plane elastic properties of impact damaged regions in composite laminates. Quasi-isotropic carbon fibre/epoxy laminates were impacted and the impact damage examined by ultrasonic C-scanning, optical microscopy and thermal deplying. After impact damage observations, specimens were cut from the laminates and tested in tension and compression. The elastic modulus of the impact damage was, in both tension and compression, mainly controlled by the amount of fibre breakage. Interestingly, layers with broken fibres could sustain some load in compression, which led to higher elastic modulus in compression than in tension. The effect of delaminations on the elastic modulus was quite small in both tension and compression. The through-the-thickness variation of in-plane stiffness was studied by successively removing plies. The variation in stiffness was negligible, probably as a result of the very uniform distribution of delaminations and fibre breakage through the thickness of the laminates.     

Impact damage resistance of CFRP with nanoclay-filled epoxy matrix

The influence of nanoclay on the impact damage resistance of carbon fibre–epoxy composites has been investigated using the low-velocity impact and compression after impact (CAI) tests. The load–energy vs. time relations were analyzed to gain insight into the damage behaviours of the materials. The CFRPs containing organoclay brought about significant improvement in impact damage resistance and damage tolerance in the form of smaller damage area, higher residual strength and higher threshold energy level. The presence of nanoclay in the epoxy matrix induced the transition of failure mechanisms of CFRP laminates during the CAI test, from the brittle buckling mode to more ductile, multi-layer delamination mode. Addition of 3 wt% clay was shown to be an optimal content for the highest damage resistance.     

复合材料层板静压入破坏机制

建立一个有效的计算模型, 以分析复合材料层板在静压入过程中发生分层、 纤维断裂的现象。该计算模型基于有限元程序的三维逐渐损伤理论对层板的静压入全过程进行模拟, 对逐层逐个单元的损伤进行判断, 可以模拟任意角度、 铺层厚度的层板在递增载荷下的逐渐损伤破坏过程。对炭纤维增强环氧树脂基复合材料层板在静压入过程中发生的分层和纤维断裂现象进行预测,并与实验结果进行比较; 对炭纤维增强双马来酰亚胺树脂基复合材料层板在静压入过程中的分层损伤和最终破坏接触力的大小进行预测,并与低速冲击下的结果进行比较。数值仿真与实验结果吻合较好, 表明静压入分析方法是复合材料层板在低速冲击下产生损伤的可替换分析方法。      

Overlay patch repair of scratch damage in carbon fiber/epoxy laminated composites

In this paper we present a numerical and experimental study on the overlay repair of scratch damage in carbon-fiber/epoxy composite laminates. The scratch damage severs several load bearing plies and results in a lack of symmetry in an originally symmetric multidirectional laminate. The ply-by-ply p-version finite element model is used to investigate the effects of the repair patch variables on the overall efficiency of the repair procedure and the lamina level stress states. The results show that interlaminar crack propagation in the direction parallel to the surface can be retarded with careful selection of repair parameters.