Abstract
An effective initial fatigue quality assessment method is presented in order to verify aircraft wing panel fastener hole whether to satisfy the design requirements. Firstly, after finishing fatigue test of bolted specimens and fatigue fracture interpretation, the time to crack initiation distributions under 3 stress levels are obtained and then a general equivalent initial flaw size distribution is established. Secondly, a method of fatigue life prediction with 95% reliability is proposed. Finally, the initial fatigue quality of aircraft wing panel fastener hole is evaluated based on the economic life criterion and double 95% EIFS value. The results show that the initial fatigue quality of the given aircraft wing panel fastener hole satisfies the design requirements.
Highlights
The results show that the initial fatigue quality of the given aircraft wing panel fastener hole satisfies the design requirements
将建立的紧固孔通用 EIFS 分布参数 α 与 Qβ 代 入公式(16) ,得到 a(0) 5/ 95 = 0.035 5 mm,满足(17) 式,也表明紧固孔原始疲劳质量满足设计要求。
Summary
摘 要:为验证某飞机机翼下壁板紧固孔细节原始疲劳质量是否满足设计要求,建立了评估其原始疲 劳质量的有效方法。 首先完成了螺栓连接试验件的疲劳试验及疲劳断口判读,建立了 3 种不同应力 水平下的裂纹萌生时间分布,并在此基础上获得通用当量初始缺陷尺寸分布;提出一种可靠度为 95% 的细节疲劳寿命预测方法,基于经济寿命准则,实现对紧固孔细节的原始疲劳质量评估;最后使用此 方法与基于满足双 95%要求的 EIFS 值对飞机机翼下壁板紧固孔原始疲劳质量进行了评估,结果表 明,该飞机机翼下壁板紧固孔细节原始疲劳质量满足设计要求。 为了模拟机翼下壁板的螺栓连接件,试验采用 螺栓连接的双狗骨型钉传载荷试件,材料为 7475⁃ T7351 铝合金,试件型式及尺寸如图 1 所示,图中尺 寸单位为 mm。 纹条带,试验中在原谱末端添加 500 次循环常幅应 力(26MPa,9MPa) 做为标识载荷。 试验载荷谱谱型 如图 2 所示。
Sign-in/Register to view highlights & summary Sign-in/Register to access full text options 
Free) 
Talk to us
Join us for a 30 min session where you can share your feedback and ask us any queries you have
Schedule a call
