Accounting for damage tolerance (DT) is crucial during the design process of aerospace composite structures. Typically, a DT design allowable limits the permitted strain level. Calculating this design allowable requires an assessment of the expected damage, the damage detectability, and the residual strength. Current state-of-the-art methods rely on empirical data, offering little flexibility and constraining the design space for a structural optimization. For a tailored calculation of the design allowable for arbitrary laminates and materials, we present an analytical analysis chain, composed from existing methods for the assessment of accidental damage (impact), the damage detectability, and the residual strength. We employ the assembled process in three steps. The determination of a DT design point for a given laminate, the calculation of a laminate-specific allowable, and the obtainment of a ply-share specific and general design allowables through minimization procedures. The highest allowables are achieved by stacking ${[-45,90,45]}_n$-blocks in the outermost plies while moving the $0\circ$-plies to the laminate center. Compared to a standard legacy quad configuration, an allowable increase between approx. $30\text{\%}$ and $50\text{\%}$ was identified. Applied in a structural optimization procedure for a composite wing, this corresponds a mass reduction of up to $5\text{\%}$.

在航空航天复合材料结构的设计过程中，考虑损伤容限（DT）至关重要。通常，DT 设计允许限制允许的应变水平。计算该设计允许值需要评估预期损坏、损坏可检测性和剩余强度。当前最先进的方法依赖于经验数据，灵活性很小，并且限制了结构优化的设计空间。为了对允许任意层压板和材料的设计进行定制计算，我们提出了一个分析分析链，由用于评估意外损坏（冲击）、损坏可检测性和残余强度的现有方法组成。我们分三步采用组装过程。确定给定层压板的 DT 设计点、计算层压板特定许用值以及通过最小化程序获得层合板特定和一般设计许用值。通过在最外层堆叠 ${[-45,90,45]}_n$ 块，同时将 $0\circ$ 层移动到层压板中心，可以实现最高允许值。与标准的传统四路配置相比，允许的增加约为。 $30\text{\%}$ 和 $50\text{\%}$ 已被识别。应用于复合材料机翼的结构优化过程中，这对应于高达 $5\text{\%}$ 的质量减少。