飞机结构广布疲劳损伤是目前大型客机损伤容限设计与分析的难点。通过试验研究了典型多孔多裂纹2024-T3铝合金平板的裂纹扩展行为。试验结果表明：相邻孔边裂纹之间的相互干扰明显降低了共线多裂纹平板的疲劳裂纹扩展寿命。就本文研究的典型多孔板，所有孔边都出现了等长裂纹这一极端情况，其裂纹扩展寿命是单孔平板孔边裂纹扩展寿命的10%左右。本文采用Eshelby夹杂理论和权函数法给出了典型多孔多裂纹问题的应力强度因子近似解析解，并结合Paris裂纹扩展公式预测疲劳裂纹扩展寿命。与采用有限元法获得应力强度因子并预测多孔多裂纹板的疲劳裂纹扩展寿命进行对比，对比结果表明：采用解析解和有限元解获得的应力强度因子预测的疲劳裂纹扩展寿命与试验结果吻合良好；相比于有限元法，本文的应力强度因子解法简单、高效，将有助于飞机结构多位置损伤（MSD）的疲劳裂纹扩展寿命预测分析。
Widespread fatigue damage poses a serious challenge to the design of modern civil aircraft. Tests are conducted to study the fatigue crack growth of 2024-T3 aluminum sheets with multiple collinear holes and cracks. It is observed that due to the interaction between adjacent hole-edge cracks, the fatigue crack growth life of the sheets with multiple collinear cracks is significantly reduced. For the extreme case in this paper, where equal-length cracks emanate from all the holes, the corresponding fatigue crack growth life is about 10% of that of two equal length cracks emanating from a single hole in a finite sheet. The Eshelby's inclusion theory and the weight function method are used to obtain the stress intensity factor of the sheet with multiple collinear holes and cracks. The approximately analytical stress intensity factors are combined with Paris's law for predicting the fatigue crack growth life. For comparison purpose, the finite element method is also used to obtain the stress intensity factor for predicting the fatigue crack growth life. Both the fatigue crack growth lives predicted by using the analytical method and finite element method agree well with the test results. Compared with the finite element method for fatigue crack growth prediction, the method proposed is simpler and more efficient, thus providing a useful tool for fatigue crack growth life prediction of the structure with Multiple Site Damages (MSD).
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