【作者】 方一帆；

【导师】 刘文博；

【作者基本信息】 哈尔滨工业大学 ， 材料学， 2010， 硕士

【摘要】 分层损伤是复合材料层合板结构主要的损伤形式,含分层损伤层合板在外载荷的作用下可能引发屈曲和分层扩展,导致层合结构在远低于设计值时发生结构破坏,这限制了复合材料结构在飞机主承力结构上的应用。对含分层损伤复合材料层合板结构的屈曲和分层扩展行为问题进行系统的研究具有十分重要的工程意义。首先通过复合材料层合板制备工艺的优化、分层缺陷定位板和封闭模具的设计制备出了符合要求的含分层损伤复合材料层合板。其次,通过试验和有限元模拟的手段对含分层损伤层合板的屈曲行为进行了研究,发现:在对复合材料层合结构的屈曲和分层特性进行研究时,采用层间特性较好的复合材料体单元,并利用无厚度的界面单元来模拟界面相的存在,可以使模拟结果和试验值有较好的吻合度。分层损伤尺寸越大,层合板的临界屈曲载荷越小。还得到了含分层损伤层合板屈曲模式:随着载荷的增加,层合板首先发生子板的变形,也即局部屈曲开始;当载荷达到整体屈曲的临界载荷时,主板位移显著增加,此时层合板开始发生整体屈曲,最终导致层合板的屈曲破坏。含不同形状分层层合板的屈曲模式是相同的;不同形状但面积相同分层层合板的屈曲临界载荷值仅存在5%左右的差别。通过对含分层损伤层合板分层扩展行为的研究发现:分层损伤尺寸越大,层合板的分层扩展临界载荷越小。但分层扩展方向不受分层尺寸的影响;分层扩展形状都是按照和初始分层相似的形状进行扩展。分层损伤形状对复合材料层合板分层扩展方向没有影响;不同形状但面积相同分层层合板的屈曲临界载荷值仅存在4%左右的差别,说明分层扩展临界载荷值对分层形状并不敏感。探讨了分层扩展与屈曲模式之间的关系:在压缩载荷下,含分层损伤层合板首先达到局部屈曲载荷发生局部屈曲,此时分层还没有发生扩展。分层扩展发生在层合板整体屈曲之前,分层扩展临界载荷略小于整体屈曲载荷,扩展方向为垂直载荷方向。当分层前缘扩展至层合板边缘时,分层扩展速度突然加快,迅速地发生贯穿,此时层合板发生失稳扩展,载荷开始下降,分层沿载荷方向以很快的速度扩展并最终导致层合板的屈曲破坏。通过上述研究为复合材料层合结构设计提供一定的理论依据,为复合材料在航天航空领域的应用提供了基础支持。更多还原

【Abstract】 It is well known that composite laminates with delamination damage is common and can cause buckling and delaminnation growth under the load,then the compressive strength of laminates will drastically reduce.This limits the application of composite laminated structures in aeronautic critical structures.So the buckling behavior and delamination growth of composite laminates containing with delamination damage must be studied.At first, through the optimization of the composite laminates process, design the delamination orientation plate and close mold,the laminates containing delamination have been successfully prepared. Then the buckling behavior and delamination growth of composite laminates containing different type of delamination under axial compressive load has been investigated by experiment and numerical analysis,the relationship between buckling mode and delamination growth is discussed.Then the buckling behavior and delamination growth of composite laminates containing different type of delamination under axial compressive load has been investigated by experiment and numerical analysis. The result shows: Using the volume cell which has good interfacial ability and interfacial cell with non-thickness simulates the buckling behavior shows good agreement with the experiment. The buckling critical load of the specimen decline as the increasing of the delaminated area. The buckling mode of composite laminates with delamination is:at first the thinner laminate starts to buckle indicates the local buckling begin. When the thinner laminate reachs the maximum out-of-plane displacement, the thicker starts to buckle, it means the global bucking of the laminates begin. The delamination shape does not affect the buckling mode. There is only 5% in the buckling critical load of laminates with different shape but the same area delamination.The delamination growth of composite laminates containing different type of delamination under axial compressive load has been study: The delamination critical load of the specimen decline as the increasing of the delaminated area. The delaminated area and shape will not affect the delaminated direction and delaminated rate. The delaminated shape is similar with the initial delaminated shape. There is only 4% in the delamination critical load of laminates with different shape but the same area delaminationThere is important relationship between delamination growth and buckling mode: when the local buckling of the laminates begins, the delamination does not spread; the delamination growth begins before the global buckling behavior. The delamination spreads along perpendicular direction of load. When the delamination spreads to the edge of the laminates, the velocity of delamination growth speed up and the direction of delamination spread turn to the load direction, the laminates soon broken.The investigations in this paper will have important improvements for composites to be more widely used in aeronautic fields.更多还原