The present study evaluates and investigates the fatigue design and fatigue life as the crack propagation occurs around the fastener of the mainframe of the central fuselage (CTR) of a rotary-wing aircraft designed based on the conventional safe-life method. For this work, the static and dynamic strain data were acquired according to flight conditions and fastening conditions between the mainframe and rail fitting (bracket). Furthermore, the fatigue life of the mainframe of rotary-wing aircraft was overestimated because of the conventional stress equation, for generating the stress spectrum according to the maneuvering conditions. To improve the existing stress equation, the modified stress equation was proposed by taking into account all force components transferred from the operating rotor system. Based on this equation, by employing the modified Goodman equation and Miner’s rule, the fatigue life of the mainframe of the CTR fuselage is calculated by the S–N diagram. Thereafter, the predicted fatigue life by considering simultaneously the damage of low-cycle fatigue and high-cycle fatigue according to the various flight conditions were compared with the actual life of crack occurrence. As a result, it showed that the predicted fatigue life (655 FHrs) was in a good agreement with the actual life (656.7 FHrs) of crack occurrence using the modified stress equation, respectively.

本研究评估和研究了疲劳设计和疲劳寿命，因为裂纹扩展发生在基于传统安全寿命方法设计的旋翼飞机中央机身 (CTR) 主机架的紧固件周围。对于这项工作，根据飞行条件和主机与导轨配件（支架）之间的紧固条件获取静态和动态应变数据。此外，由于传统的应力方程，旋转翼飞机主机架的疲劳寿命被高估，以根据机动条件生成应力谱。为了改进现有的应力方程，通过考虑从运行的转子系统传递的所有力分量，提出了修正的应力方程。基于这个方程，采用修正的古德曼方程和米纳法则，通过S-N图计算出CTR机身主机架的疲劳寿命。此后，根据各种飞行条件同时考虑低周疲劳和高周疲劳的损伤，预测的疲劳寿命与裂纹发生的实际寿命进行了比较。结果表明，分别使用修正的应力方程预测的疲劳寿命（655 FHrs）与裂纹发生的实际寿命（656.7 FHrs）非常吻合。根据不同飞行条件同时考虑低周疲劳和高周疲劳的损伤，预测的疲劳寿命与裂纹发生的实际寿命进行了比较。结果表明，分别使用修正的应力方程预测的疲劳寿命（655 FHrs）与裂纹发生的实际寿命（656.7 FHrs）非常吻合。根据不同飞行条件同时考虑低周疲劳和高周疲劳的损伤，预测的疲劳寿命与裂纹发生的实际寿命进行了比较。结果表明，分别使用修正的应力方程预测的疲劳寿命（655 FHrs）与裂纹发生的实际寿命（656.7 FHrs）非常吻合。