层压板修理设计中的参数选择问题

复合材料层压板的修理问题是复合材料结构修理中的基础性研究课题。其中,贴补法和挖补法是两种基本的修理方法。通过有限元方法对贴补法和挖补法进行了研究,给出了贴补修理和挖补修理的最佳设计参数选择方案。对于大多数贴补修理设计,补片的厚度应该是母板厚度的一半,搭接长度大约在25mm;对于挖补修理,通常可采用6°的挖补角度。试验结果表明,通过计算分析得出的最佳修理设计参数对工程实际具有很好的指导意义。      

复合材料层压板单面贴补修理建模分析研究

胶接修理技术是一种优质、高效、低成本的结构修理技术.针对复合材料层板的胶接贴补修理进行计算模拟分析,并与实验结果相对比.采用ABAQUS有限元软件对层压板的损伤板和单面贴补修理板进行有限元建模,分析修补前后的应力分布情况,计算出失效强度与完好板进行对比,得出修补后的强度恢复率.     

干法/湿法挖补修理层合板力学性能的对比研究

挖补修理是一种常用的受损复合材料结构修复方法,主要可分为干法(预固化补片)修理和湿法(湿铺贴)修理。通过开展静力拉伸实验,研究了不同损伤尺寸的预固化补片挖补修理和湿铺贴挖补修理层合板的极限强度、破坏模式以及补片脱粘情况,实验结果表明采用预固化补片修理的结构其强度恢复率更高,补片提前脱粘概率更低,修理效果更好。在力学实验的基础上,通过有限元计算准确预测了挖补修理层合板的极限强度、应力应变分布和损伤演化过程,计算结果表明预固化补片修理结构的胶层应力分布更加均匀,应力水平和应力集中程度更低,出现界面失效和胶层内聚失效的载荷也更高。     

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。     

层压板双面挖补修理的拉伸性能研究及参数分析

对含穿透型损伤层压板双面挖补胶接修补件的拉伸性能及主要影响参数进行了试验研究。结果表明: 双面挖补后的失效拉伸强度恢复率能够达到80%; 当挖补斜度1∶40、 覆盖层取3层、 且靠近母板的铺层方向与母板的最外层的铺层方向保持一致时, 强度恢复率较大; 使用双面挖补后, 母板厚度、 母板铺层方式对强度恢复率的影响不大。同时应用有限元软件ABAQUS, 采用三维实体单层设置建模方式对双面挖补层压板的拉伸性能进行了数值模拟, 模拟计算结果与试验结果基本相符。      

挖补复合材料层合板的拉伸和压缩性能研究现状EI北大核心CSCD

随着复合材料层合板结构应用的不断扩大,其修理问题也日益凸出。挖补修理在层合板修理中占有举足轻重的地位,而挖补后结构的拉伸和压缩性能恢复是表征修理质量极其重要的指标,因而对挖补复合材料层合板拉伸和压缩性能的研究具有重要意义。文中总结了挖补层合板拉伸和压缩性能的研究现状,对材料、工艺、构型及环境等参数的影响进行了分析,据此给出了挖补修理优选方案,可供复合材料结构挖补修理设计参考。     

蜂窝夹层修理结构的弯曲性能试验分析

采用四点弯加载方式研究分析了含损伤蜂窝夹层修理结构的弯曲性能,该夹层结构由碳纤维增强的聚合物面板和蜂窝芯子组成。进一步分析了挖补斜度、挖补方式、损伤程度、修理设备和修理材料对修理板弯曲性能的影响。研究表明,修理板的破坏模式可分为补片边缘折断、补片中面折断和胶层破坏三种,相同破坏模式修理板的名义弯曲强度相近,其中前两种破坏模式修理板的名义弯曲强度与完好板相近,而第三种破坏模式修理板的名义弯曲强度相对较低。所有修理板的名义弯曲强度恢复率基本处于95%以上,同时修理后抗弯刚度也满足修理准则。     

挖补修理复合材料层合板疲劳性能研究进展

随着纤维增强复合材料结构的应用更加广泛,其结构修理问题日渐凸显。挖补修理在复合材料层合板修理中具有重要地位,而修理后结构的疲劳性能是评价结构修理质量的重要依据,因此对复合材料挖补修理层合板疲劳性能的研究意义重大。文中从模型和试验两方面总结了挖补修理复合材料层合板疲劳性能研究现状,指出了当前挖补修理复合材料层合板疲劳性能研究存在的问题,可为复合材料结构挖补修理研究和工程应用提供参考。     

飞机复合材料结构修理设计渐进损伤研究

复合材料在现代飞机结构中所占比重越来越大。从现阶段大量的金属飞机结构繁重的维修任务可以看出,未来复合材料飞机结构维修任务的艰巨性。复合材料结构修理渐进损伤研究已经成为未来复合材料技术研究的重要方向。本文从复合材料层合板贴补修理分析与强度预测分析入手,讨论复合材料层合板挖补修理技术。     

复合材料层合板胶接贴补修理渐进损伤分析

建立了复合材料层合板胶接贴补修理构型渐进损伤分析的三维有限元模型, 其中层合板和胶层分别采用正交各向异性损伤和各向同性损伤的连续介质损伤力学模型, 整个分析过程中同时考虑层合板和胶层的损伤形成和扩展以及它们之间的相互影响, 单向压缩载荷作用下的层合板贴补修理构型的试验数据验证了该模型的有效性, 采用该模型分析了不同的贴补修理参数对修补强度的影响。 结果表明: 当层合板补片较薄时, 补片损伤是导致修补结构失效的主要原因; 当补片较厚时, 胶层失效是导致修补结构失效的主要原因, 此时补片厚度增加并不能显著增大修补结构的极限强度。在复合材料贴补修理时需要对补片和胶层进行详细优化设计。      

含半穿透损伤层合板挖补修理后的拉伸性能

对含半穿透损伤层板挖补修理后的拉伸性能进行了试验研究, 结果表明修理试件的拉伸强度和破坏模式随挖补斜度的变化出现显著差异。对修理试件的拉伸性能进行了有限元模拟, 计算得到的极限强度和破坏模式与试验结果吻合良好。数值模型计算结果表明, 挖补斜度是修理试件最重要的设计参数, 其对试件的极限强度、 破坏模式及修理/未修理子层间界面损伤均有显著影响。研究结论可以为含半穿透损伤层板的挖补修理设计提供理论指导。     

单面贴补修理后层合板的拉伸性能

对含孔损伤复合材料层合板单面贴补后进行拉伸试验研究。测量了层合板的应变分布、修理后层合板中心点的离面位移及拉伸强度等, 考察了补片的厚度、大小等因素对修理效果的影响。结果显示, 增加补片的厚度和直径能够提高母板的承拉能力, 但是增加补片的厚度会导致层合板离面位移增大。对无侧边支持的单面贴补层合板进行计算分析时, 必须考虑偏心载荷引起的弯矩的影响。在此基础上, 采用分层损伤判据建立了三维有限元模型, 对单面贴补层合板的破坏机理和拉伸强度进行了计算和分析。结果表明, 修理后层合板的拉伸破坏是由补片或母板内与胶接面相邻的层间分层引起的; 计算结果与试验结果一致。      

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.     

Optimum shapes of scarf repairs

Adhesively bonded scarf repairs are the preferred method for repairing composite structures, limited mainly by the amount of material removal associated with scarfing. In addition to the high strength restoration, scarf repairs also enable recovery of the original external surface as required by aerodynamic and/or external mould line considerations. However, scarf repairs almost inevitably result in the removal of undamaged material to make way for the scarf insert. This can be a particularly significant issue for thick structures, because the scarf length can vary between 20 and 100 times the thicknesses of the parent structure. In this investigation, an optimisation method has been developed for determining the optimum repair shapes for a given biaxial loading condition. The optimum scarf shape is determined by numerically solving the resulting non-linear differential equation governing the scarf angle. The optimum and near-optimum shapes are presented and discussed with respect to computational modelling using the finite element method.     

Failure load prediction of laminates repaired with a scarf-bonded patch using the damage zone method

The damage zone method (DZM) is an efficient way to predict the failure of composite structures with a minimum of real testing. Particularly, it is useful when the failure mechanism is too complicated to be accurately analyzed by a merely numerical method. The aim of this study was to use the damage zone model to predict the failure load of repaired laminates, in which scarf-bonded joints were used for repair. The model uses a test-based critical damage zone and stress-based failure criteria. A total of 45 carbon-epoxy composite (USN) laminate scarf-repaired specimens were first tested with two different defect sizes, four scarf angles, and three overlap layer sizes. The Tsai-Wu and Tsai-Hill criteria were used for the laminate, and the maximum shear stress criterion for the adhesive was adopted to predict failure onset. The predicted failure loads were compared to test results and a good agreement was obtained with a 9.2% maximum deviation for almost all parameters with the exception of a case with an unrealistically large scarf angle. To verify the feasibility of the DZM for different material, additional 30 repair specimens using other unidirectional carbon-epoxy laminate were then also tested and the predictions were confirmed by the results of the experiment.     

Evaluation of fatigue behavior on repaired carbon fiber/epoxy composites

The continuous use of structural polymer composites in aeronautical industry has required the development of repairing techniques of damages found in different types of laminates. The most usually adopted procedure to investigate the repair of composite laminates has been by repairing damages simulated in laminated composite specimens. This work shows the influence of structural repair technique on mechanical properties of a typical carbon fiber/epoxy laminate used in aerospace industry. When analyzed by tensile test, the laminates with and without repair present tensile strength values of 670 and 892 MPa, respectively, and tensile modulus of 53.0 and 67.2 GPa, respectively. By this result, it is possible to observe a decrease of the measured mechanical properties of the repaired composites. When submitted to fatigue test, it is observed that in loads higher than 250 MPa, this laminate presents a low life cycle (lower than 400,000 cycles). The fatigue performance of both laminates is comparable, but the non-repaired laminate presented higher tensile and fatigue resistance when compared with the repaired laminate.     

Battle damage repair of a helicopter composite frame-to-skin junction – A sole external repair approach

Research was conducted to examine the effectiveness of a rapid repair to a helicopter composite frame-to-skin junction subjected to battlefield damage. The repair design consists of a laminate patch and aluminium angle bracket adhesively bonded and riveted, respectively, to the helicopter external surface. The assessment involved a relative comparison of three models, representing pristine, damaged and repaired configurations. Computational analyses were conducted to examine the stiffness and buckling onset load of the overall structure and the strengths of individual components (laminates, adhesive bondlines and rivets) under three typical load conditions, namely in-plane shear, axial compression and transverse compression. The results showed that the damage would cause significant stiffness and strength reduction. The repair could sufficiently restore the stiffness and static strength for the load cases considered. However, for the specimen without support from its adjacent helicopter structure, it is predicted that the failure mode under the transverse compression loading would be via buckling under a relatively low load. A compression test was conducted to further validate the repair design. The result agreed well with the prediction. It showed that compared with an un-repaired damaged specimen, the external repair increased the strength by 83%. The equivalent far field failure strain exceeded 3300 με which is considered satisfactory for a rapid field battle damage repair (BDR).     

Effect of Boundary Condition on Impact Response and Damage Meso-mechanism for Scarf Bonded Repair of Load-bearing Composite Structures

Composite scarf repair is applied in airplane load-bearing structures. But the repairs would have different boundary supports in real situation,such as simply support,clamped support and flexible support. With different boundary supports, the bonded scarf repair of composite structures could have diverse damage mechanisms and responses. This work intends to study the impact performance of different boundaries. For this purpose,experimental tests were carried out on the specimens with two sides support and four sides support. The impact load and deflection was monitored during the tests. After impact being finished,the meso-mechanism was studied by means of micro-cracks of the side section for two sides support condition and cross section for the four sides support condition. It was possible to conclude that the four sides boundary possess higher impact resistance maximum loads,lower displacements and lower extent load dropping. In terms of damage modes,as the impact energy relative low 12 J and 16 J,the central position of scarf bonding zone for two sides support appears adhesive cohesive failure and adhesive-composite interface failure. When the energy increases to 20 J,the dominated damage of two sides support moves down to the feathered tip. For four sides support,the critical energy level is 25 J,under which the scarf adhesive begins to be damaged.