Fatigue Crack Growth in Complex Residual Stress Fields due to Surface Treatment and Foreign Object Damage under Simulated Flight Cycles

航空发动机关键部件经常发生高周疲劳(HCF)和低周疲劳(LCF)的联合作用，损伤容限方法已被应用于其结构完整性评估。低周期与起飞和着陆之间的周期相同，而高周期是飞行中空气动力学引起的振动的结果。航空发动机部件在高周疲劳和低周疲劳联合加载条件下进行疲劳完整性评估，以模拟服役加载条件下的疲劳完整性是势在必行的。30多年来，朴茨茅斯团队与劳斯莱斯、QinetiQ和USAir Force密切合作，一直是研究HCF和LCF组合加载条件下疲劳裂纹扩展的领先指标之一。近年来，异物损伤已被认为是汽轮机叶片的主要寿命限制因素之一。在起飞和着陆过程中，由于吸入微小的硬颗粒而产生的撞击可达到500m/S的速度，并对航空发动机的翼型造成严重破坏。据估计，FOD每年给航空发动机行业造成的损失为40亿美元。弹丸撞击引起的疲劳强度降低仅在HCF加载条件下进行过研究。在朴茨茅斯进行了第一次关于FOD下组合的HCF和LCF对裂纹扩展的影响的研究(GR/R79258，提交的最终报告)。在残余应力及其对裂纹驱动力的影响方面取得了重大进展。具体而言，在FOD形成的凹坑根部附近存在拉应力，拉应力深度可达0.2 mm以上。在施加HCF和LCF加载块时，局部应力比增加，促使早期裂纹优先从这些位置扩展，导致裂纹扩展速率显著增加，与单独使用HCF或LCF加载相比。对于翼型，只有在计算裂纹驱动力时考虑FOD引起的残余应力，才能揭示裂纹的加速扩展。翼型的前缘特别容易受到FOD的影响。近年来，激光冲击喷丸(LSP)或低塑性抛光(LPB)等表面处理方法的引入，显著提高了材料的疲劳强度和抗裂纹扩展能力。这种处理的目的是在风扇叶片的前缘产生显著的压缩残余应力，从而使叶片的临界区域变得更耐损坏，以避免灾难性故障。通常，使用LSP或LPB可以获得1-2 mm的压缩残余应力深度，而不是传统的喷丸强化约0.2 mm。为了最大限度地提高断裂关键部件的疲劳抗力，在LSP/LSB存在的情况下，对FOD引起的疲劳损伤过程有一个基本的了解是至关重要的。该模型将考虑表面处理和FOD引起的残余应力的影响。将对动态冲击进行模拟，并利用有限元分析和X射线衍射法对残余应力的稳定化和松弛进行研究。将对表面处理的有效性进行评估，并使用模拟飞行试验数据对模型进行验证。如果要准确地模拟FOD引起的裂纹起始和早期扩展，这样的研究是至关重要的，这样就可以为航空发动机工业提供FOD影响的疲劳完整性的预测工具。这项工作有助于航空发动机关键部件(如涡轮叶片)的安全设计和寿命管理。

Damage tolerance approach has been employed in the assessment of structural integrity of critical aeroengine components, where a conjoint action of high cycle fatigue (HCF) and low cycle fatigue (LCF) often occurs. The low cycles are identified with the period between takeoff and landing, while high cycles are a result of inflight aerodynamically induced vibrations. It is imperative that fatigue integrity assessment of aero engine components is carried out under the combined HCF and LCF loading conditions to simulate in service loading conditions. The Portsmouth team has been one of the leading exponents in the study of fatigue crack growth under combined HCF and LCF loading conditions, in close collaboration with Rolls-Royce, QinetiQ and USAir Force for over 30 years. In recent years, foreign object damage (FOD) has been identified as one of the main life limiting factors for turbine blades. Impacts due to small hard particle ingestion during takeoff and landing can reach velocities up to 500 m/s and cause severe damage to aerofoils in aero engines. Damage due to FOD is estimated at 4 billion US dollars annually for the aeroengine industry. Reduction in fatigue strength due to projectile impacts has been studied exclusively for HCF loading conditions. The first study on the effects of combined HCF and LCF under FOD on crack growth was carried out at Portsmouth (GR/R79258, final report submitted). Significant progress has been made in the assessment of residual stresses and their effect on crack driving force. Specifically, tensile residual stresses were found in the vicinity of the crater root made by FOD; and the depth of these tensile stresses can reach to more than 0.2 mm. Upon application of a combined HCF and LCF loading block, the local stress ratios were elevated which prompted early crack growth preferentially from these sites, resulting in significantly increased crack growth rates, as compared with either HCF or LCF loading alone. For aerofoils, the accelerated crack growth was only revealed when the residual stresses due to FOD were considered in the calculation of crack driving force. The leading edge of aerofoils is particularly susceptible to FOD. In recent years, the introduction of surface treatments, such as laser shock peening (LSP) or low plasticity burnishing (LPB) have significantly improved the fatigue strength and crack growth resistance. Such treatments aim at producing significant compressive residual stresses along the leading edge of fan blades, such that the critical region of the blades becomes considerably more damage tolerant to avoid catastrophic failures. Typically, the depths of the compressive residual stresses can be achieved using LSP or LPB are 1-2 mm, as opposed to ~0.2 mm by conventional shot peening. A fundamental understanding of fatigue damage process due to FOD in the presence of LSP/LSB is vital, if they are to be utilized to the full potential in enhancing fatigue resistance of fracture critical components, particularly in the event of FOD.The proposed research aims at developing a predictive model for fatigue crack onset and early growth under simulated flight cycles. The model will take into account of the effect of residual stresses due to surface treatment as well as FOD. Dynamic impact will be modelled and the stabilised and the relaxation of residual stresses will be studied using the finite element analysis and validated by the X-ray diffraction method. The effectiveness of the surface treatment will be evaluated and the model will be validated using simulated inflight test data. Such studies are critical if onset and early crack growth due to FOD is to be modelled accurately, so that predictive tools may be made available to aeroengine industry for FOD-affected fatigue integrity. The work contributes to the safe design and life management of critical aeroengine components such as turbine blades.