孔边倒角和预腐蚀作用下航空铝合金疲劳性能及断裂机理研究

基于航空铝合金带孔结构材料在服役过程中常因腐蚀损伤而导致疲劳断裂问题,通过对未腐蚀和预腐蚀24h后的7075铝合金双孔未倒角和双孔倒角试样进行疲劳实验研究,分析腐蚀预损伤和孔边倒角对试件疲劳性能的影响及疲劳断裂特性差异。结果表明:腐蚀预损伤对7075铝合金材料疲劳寿命的影响显著,双孔未倒角和倒角试样预腐蚀24h后试样中值疲劳寿命比未腐蚀试样最大下降了31.74%和26.92%;孔边倒角对材料疲劳寿命有一定的影响,未腐蚀和预腐蚀24h孔边倒角试样中值疲劳寿命比未倒角试样最大下降了28.02%和15.36%,主要原因是由于孔边倒角过程中产生加工刀痕,引入了"预损伤",且倒角后疲劳裂纹萌生位置变多,导致材料发生疲劳断裂的概率变大。     

腐蚀环境对预腐蚀铝合金腐蚀疲劳性能的影响

采用不同环境下的轴向加载疲劳性能测试方法,研究了不同环境对预腐蚀7XXX铝合金腐蚀疲劳性能的影响.结果表明:腐蚀环境对其疲劳寿命的影响从重到轻依次是:油箱积水、盐水、潮湿空气和实验室空气.当应力比R为0.5和0.06时,潮湿空气下的疲劳寿命比实验室空气下的略低,降低的幅度随着应力的降低而增大;在低应力水平区盐水和油箱积水环境对疲劳寿命的影响基本相同.当应力比R为-1时,盐水下的疲劳寿命总是略低于油箱积水环境下的疲劳寿命;潮湿空气相比实验室空气下的疲劳寿命随着应力的降低,降低的幅度比应力比为0.5和0.06时要略大.     

2A12铝合金疲劳裂纹的成核与扩展机理EI北大核心CSCD

进行了不同应力水平下和不同应力比下2A12铝合金试验件的疲劳试验,并对试验件疲劳断口显微结构进行分析和对比,以揭示疲劳裂纹成核与扩展的微观特征。结果表明:2A12铝合金的疲劳裂纹在靠近试验件表面较为粗大的第二相粒子或试验件棱角缺陷处成核;成核位置距试验件表面的距离与应力水平和应力比有关,疲劳裂纹扩展区域也与应力水平和应力比有关。应力水平较小或应力较大时,疲劳裂纹沿晶扩展和穿晶扩展尤为明显。疲劳断口的裂纹成核粒子越小(大),试验件疲劳寿命越短(长),微裂纹成核寿命是疲劳全寿命计算中不可忽视的部分。     

LC4铝合金元件的疲劳寿命分析EI

用扫描电镜和光学金相研究了ＬＣ４铝合金疲劳试件的宏、微观断裂特征，论述了试件断裂位置转移和疲劳寿命分散的原因，讨论了加工与装配质量对疲劳寿命的影响。     

LC4CS高强铝合金腐蚀疲劳行为研究EI

本文较全面地研究了室温空气、潮湿空气、盐雾、盐雾加二氧化硫、人造海水环境下的ＬＣ４ＣＳ高强铝合金抗疲劳和裂纹扩展行为，揭示了各种环境介质对ＬＣ４ＣＳ高强铝合金的影响。本文将ＤＦＲｃｕｔｏｆｆ这一新概念引入腐蚀疲劳领域，为新老飞机寿命评估提供了重要依据。     

3.5%NaCl腐蚀环境下2种航空铝合金材料疲劳性能试验研究

腐蚀环境下的疲劳性能是航空金属结构疲劳寿命设计的重要前提,为此,试验测定了2种航空铝合金材料（2E12-T3、7050-T7451）的光滑试样和缺口试样在干燥大气和3.5%NaCl腐蚀环境下的疲劳性能,在试验数据的基础上进行性能对比,并对试样断口进行扫描电镜（SEM）分析,研究了3.5%NaCl腐蚀环境与载荷联合作用对腐蚀疲劳性能的影响机理,研究结果表明：3.5%NaCl腐蚀环境对2种铝合金材料的疲劳性能均产生不利影响,且腐蚀与疲劳载荷的交互作用随着应力水平的降低而增强,疲劳性能下降更明显;与光滑试样相比,腐蚀环境对铝合金2E12-T3缺口试样疲劳性能的影响更大,但对铝合金7050-T7451缺口试样疲劳性能的影响却变小;在腐蚀环境下,裂纹尖端易发生电化学反应产生腐蚀产物和[H]离子,腐蚀产物的存在会阻碍裂纹闭合,同时,[H]离子导致裂纹尖端的氢脆效应,加快裂纹扩展,使疲劳性能降低。     

晶粒度对K492高温合金疲劳性能的影响EI北大核心CSCD

研究了晶粒度对K492镍基高温合金在700℃和800℃的疲劳性能的影响,并根据扫描电镜和透射电镜的结果分析了疲劳断裂的机理。结果表明,晶粒细化提高了K492合金700℃和800℃的疲劳性能。在700℃进行高周疲劳时在冶金缺陷或某个晶面处产生疲劳裂纹;疲劳实验后位错呈带状分布,γ'相形态不发生变化,位错以切过或Orowan绕过机制通过γ'相。在800℃高周疲劳时疲劳裂纹均在缺陷处产生,部分区域的位错组态与在700℃的实验结果相似,γ'相形态不发生变化。另一部分区域的γ'相发生筏化,位错分布在基体通道中,γ'相失去对位错的钉扎作用。在700℃低周疲劳时疲劳裂纹主要起始于样品表面。在800℃低周疲劳时,疲劳裂纹主要产生于样品次表面或某个晶面处。     

晶粒度对K492高温合金疲劳性能的影响EI北大核心CSCD

研究了晶粒度对K492镍基高温合金在700℃和800℃的疲劳性能的影响,并根据扫描电镜和透射电镜的结果分析了疲劳断裂的机理。结果表明,晶粒细化提高了K492合金700℃和800℃的疲劳性能。在700℃进行高周疲劳时在冶金缺陷或某个晶面处产生疲劳裂纹;疲劳实验后位错呈带状分布,γ'相形态不发生变化,位错以切过或Orowan绕过机制通过γ'相。在800℃高周疲劳时疲劳裂纹均在缺陷处产生,部分区域的位错组态与在700℃的实验结果相似,γ'相形态不发生变化。另一部分区域的γ'相发生筏化,位错分布在基体通道中,γ'相失去对位错的钉扎作用。在700℃低周疲劳时疲劳裂纹主要起始于样品表面。在800℃低周疲劳时,疲劳裂纹主要产生于样品次表面或某个晶面处。     

预腐蚀后拉伸超载对LC4CS铝合金疲劳性能的影响

采用预腐蚀后进行单次拉伸超载疲劳试验的方法，研究了腐蚀与超载对LC4CS材料疲劳寿命的影响。结果表明，如果保持超载比不变，试件疲劳寿命将随预腐蚀年限增加而线性降低，且分散性逐年增大；在低超载比时，预腐蚀对疲劳寿命的亏损作用与超载对疲劳寿命的增益作用相互抵消；在超载比较大时，超载迟滞效应对疲劳寿命的增益作用明显。     

油箱积水环境对预腐蚀铝合金腐蚀疲劳性能的影响

对预腐蚀2XXX铝合金Kt=1和Kt=3试样进行了实验室空气环境和油箱积水环境下的疲劳寿命实验,并用单侧容限因数法对疲劳安全寿命进行了估算,得到了预腐蚀2XXX铝合金在这两种环境下的中值疲劳寿命和99%可靠度95%置信度水平上的安全疲劳寿命。油箱积水环境显著降低了预腐蚀2XXX铝合金的疲劳性能,实验室空气环境下的疲劳寿命数据比较分散,应力水平越低越分散,而油箱积水环境下的疲劳寿命数据比较集中。油箱积水环境下的腐蚀主要由电化学腐蚀和微生物腐蚀组成。     

潮湿空气环境对2024-T3铝合金疲劳性能的影响

采用2024-T3铝合金含中心孔试件进行了实验室空气环境和潮湿空气环境下的疲劳寿命实验及升降法实验,研究了潮湿空气环境对2024-T3铝合金疲劳性能的影响.结果表明,潮湿空气环境显著降低了2024-T3铝合金的疲劳性能,潮湿空气环境对其疲劳寿命特征值和疲劳强度均值的影响系数分别为0.6059和0.8722;在潮湿空气环境中疲劳寿命的分散性变大,用疲劳寿命的中值或特征值得到的腐蚀影响系数进行可靠度95%的腐蚀环境下的寿命修正,将得到偏危险的结果.用潮湿空气环境对基本可靠性寿命的影响系数由标准S-N曲线折算得到的对疲劳强度的影响系数与升降法测得的对疲劳强度的影响系数基本一致,在潮湿空气环境下标准S-N曲线参数仍然适用.     

High cycle fatigue characteristics of 2124-T851 aluminum alloy

The fatigue crack growth rate, fracture toughness and fatigue S-N curve of 2124-T851 aluminum alloy at high cycle fatigue condition were measured and fatigue fracture process and fractography were studied using optical microscopy (OM), X-ray diffraction (XRD) technique, transmission electron microscopy (TEM) and scanning electron microscopy (SEM). The results show that at room temperature and R = 0.1 conditions, the characteristics of fatigue fracture could be observed. Under those conditions, the fatigue strength and the fracture toughness of a 2124-T851 thick plate is 243 MPa and 29.64 MPa · m^1/2, respectively. At high cycle fatigue condition, the higher the stress amplitude, the wider the space between fatigue striations, the faster the rate of fatigue crack developing and going into the intermittent fracture area, and the greater the ratio between the intermittent fracture area and the whole fracture area.     

Corrosion fatigue behaviour and fracture mechanisms in nitrided low alloy steel

Rotary bending fatigue tests have been performed in 3%NaCl aqueous solution using specimens of a low alloy steel (Cr–Mo steel) with different nitride case depths. The effect of case depth on corrosion fatigue strength, the fracture process and mechanisms were studied. The corrosion fatigue strengths of the nitrided materials increased compared with the untreated material and increased with increasing thickness of the compound layer, but tended to saturate above a certain thickness. All the materials showed lower fatigue strength in 3%NaCl aqueous solution than in laboratory air and the reduction of fatigue strength decreased with increasing thickness of the compound layer, but remained nearly constant above a certain thickness. Corrosion pits were seen underneath the compound layer, from which cracks initiated. The corrosion fatigue strengths of the specimens whose compound layer was completely removed by electropolishing were almost the same as that of the untreated material, indicating a very significant role of the compound layer in improving corrosion fatigue strength. Because of the porous nature of the compound layer, particularly in the surface‐adjacent part, the solution penetrated the compound layer and reached the base steel, thus the corrosion fatigue strength of the nitrided materials was controlled by the penetration of corrosive media.     

Intergranular corrosion of a light aircraft aluminum alloy propeller

Examination of the wreckage of a light aircraft revealed that approximately 20 cm was missing from one tip of the aluminum alloy propeller. Fractographic and metallographic examination of the remaining portion of the propeller revealed extensive grain-boundary separation in the vicinity of the fracture, and grain edges and corners rounded by corrosion on the fracture surface. Energy dispersive X-ray analysis (EDXA) revealed fluorine on, and in the vicinity of, the fracture surface. In the ensuing litigation, it was asserted that the crash occurred because the propeller fractured in flight as the result of intergranular attack caused by the use of a fluorine-bearing cleaner.     

On the corrosion fatigue crack initiation model and expression of metallic notched elements

In the present study, attempts are made to extend the application of the mechanical model for the fatigue crack initiation (FCI) and the FCI life formula of metallic notched elements in laboratory air to those in the corrosive environment. The test results and analysis of the corrosion FCI (CFCI) life of aluminum alloys and Ti---6A1---4V show that the expression of the CFCI life obtained by modifying the FCI life formula in laboratory air can give a good fit to the test results of the CFCI life. The salt water (3.5% NaCl) environment has no effects on the CFCI resistant coefficient compared with the FCI resistant coefficient in laboratory air. However, 3.5% NaCl environment greatly decreases the CFCI threshold of aluminum alloy, but has little effect on the CFCI threshold of Ti---6A1---4V. The loading frequency ranging from 1 Hz to 10 Hz has no appreciable effect on the CFCI life, and thus, the CFCI threshold of aluminum alloys investigated. Hence, the expression for the CFCI life of metallic notched elements proposed in this study is a better one, which reveals a correlation between the CFCI life and the governing parameters, such as, the geometry of the notched elements, the nominal stress range, the stress ratio, the tensile properties and the CFCI threshold. However, this new expression of the CFCI life needs to be verified by more test results.     

选区激光熔化AlSi10Mg合金不同温度原位拉伸变形行为及断裂机理研究

为了提高选区激光熔化AlSi10Mg合金在航空航天领域的应用,基于自主研发的原位SEM高温拉伸台,本文对比分析了原位拉伸非校准样品的选区激光熔化AlSi10Mg合金在室温、200、300 ℃条件下的力学性能与显微组织动态演化,并总结了断裂机理。结果表明,选区激光熔化AlSi10Mg合金的显微结构由α-Al基体、共晶Si和大量的气孔组成,且共晶Si呈连续网格状均匀分布在α-Al基体上。随着温度的升高,选区激光熔化AlSi10Mg合金的强度降低。屈服强度从室温的207 MPa降低到300 ℃时的52 MPa,极限抗拉强度从室温的304 MPa降低到300 ℃时的71 MPa,延伸率则随温度的升高而增大,从室温的7.4%增大到300 ℃时的59.5%。室温拉伸过程中试样并未出现明显的颈缩现象,而是随着温度的升高,试样的颈缩现象逐渐明显,表明试样经历了更加充分的塑性变形, 并且随着温度的升高,试样的断裂位置越来越偏离标距段中心。通过对试样变形行为的研究发现,200 ℃时,变形主要集中在晶内,发生晶内滑移;而300 ℃时滑移主要集中在晶界,导致晶界滑移。由于试样表面及内部存在大量缺陷,因此,室温下选区激光熔化AlSi10Mg合金的断裂机理为熔池边界的组织突变结合孔洞连通造成的准解理断裂。随着温度的升高,由于初始孔洞边缘的应力集中产生新的孔洞形核,新形核的孔洞与相邻孔洞相连通,导致试样的最终断裂。     

Stellite12钴基合金的疲劳性能及其断裂机理研究

采用三点弯曲疲劳法测得光滑试样和直缺口试样的S-N曲线以研究Stellite12钴基合金的疲劳性能,并通过断口形貌观察进一步探究该钴基合金的断裂过程。结果表明:光滑试样的疲劳极限为545 MPa,约为原始抗弯强度1552 MPa的25.4%;直缺口试样的疲劳极限约为101MPa,约为静态抗弯强度517.6MPa的19.1%。对于疲劳敏感性,光滑试样与直缺口试样的疲劳敏感性分别为397和31。此外发现疲劳裂纹多萌生于近表层聚集的碳化物处,同时表面缺陷也可诱发疲劳裂纹的萌生。疲劳裂纹的扩展主要表现为碳化物的穿晶断裂,钴基体在应力比R=0.1的疲劳加载条件下虽表现出一定的韧性且呈现出较多的撕裂脊,但也呈现出一定的脆性断裂模式,因此疲劳裂纹扩展模式为真疲劳与静态疲劳的混合模式。     

Investigation of fracture mechanism of 6063 aluminum alloy under different stress states

Fracture mechanisms in a 6063 aluminum alloy were investigated and analyzed carefully by in-situ tensile tests in SEM with a vacuum chamber. Specimens used were designed to produce different stress states. Studies indicated that with stress triaxiality (σ _m/σ _e) decreasing, the fracture modes changed from normal fracture to shear fracture and the fracture surfaces changed from the dimples and intragranular dominated fracture mode to the shear dominated fracture mode. The grain boundaries of the 6063 aluminum alloy were the weakest positions. In the case of high stress triaxiality, the grain boundary cracks were produced by normal stress or by the incompatibility of deformation between neighboring grains, and the normal stress dominated the crack propagation. In the case of low stress triaxiality, the boundary cracks were produced by the relative slipping of grains against neighboring grains, and the shear stress dominated the crack propagation. The final fracture of the specimens occurred by connections of cracks through transgranular cracking of the ligaments among these cracks.     

Research on corrosion performance of 6061 aluminum alloy in salt spray environment

Salt spray corrosion test was carried out on 6061 aluminum alloy, and quasi-static tensile test at room temperature was carried out on the sample with universal testing machine. The effect of salt spray corrosion on the mechanical properties of 6061 aluminum alloy was studied by scanning electron microscope (SEM), energy dispersive spectrometer (EDS) and electrochemistry. The corrosion rate of 6061 aluminum alloy was quantitatively characterized by different corrosion parameters. It was found that local corrosion of 6061 aluminum alloy occurred in salt spray environment, mainly pitting corrosion and intergranular corrosion. With the increase of corrosion time, the polarization resistance of 6061 aluminum alloy decreases, and the corrosion rate significantly increases. The average corrosion rate and the maximum corrosion rate of 6061 aluminum alloy were characterized by corrosion weight loss and corrosion pit depth. And they can be transformed into each other. The mechanical properties of 6061 aluminum alloy were mainly affected by the depth of corrosion pit. With the increase of corrosion time, the tensile strength and fracture strain decreased, resulting in poor plasticity of the sample. At the same time, the change of elongation of 6061 aluminum alloy can be accurately predicted by the depth of corrosion pit.     

Analysis and prediction of microstructural effects on long-term fatigue performance of an aluminum aerospace alloy

A program of experimental and analytical tasks has been conducted to define the linkage(s) between microstructural characteristics and fatigue performance in an aluminum alloy typically used for airframe structural applications. The first goal was to develop data for quantitatively linking measurable characteristics of material microstructure with long-term fatigue performance. The second goal was to develop models to predict fatigue performance based on the microstructural characteristics. The work focused on several process variants of aluminum alloy 7050-T7451 plate. This material was chosen because of its widespread use for flight-critical airframe structural components, and the particular characteristics associated with the manufacturing, service and maintenance of thick section components. Within the framework of this objective, life-limiting microstructural features have been identified and ranked by severity, and models to quantitatively describe the evolution and growth of macrostructural cracks from those features have been developed.The modeling framework has been applied to predict the cyclic lifetime of the 7050 alloy process variants based on the populations of life-limiting microstructural features. In addition, the models have been used to show how changes in the material characteristics may affect the fatigue performance. This includes predictions of the effect of changing the life-limiting microfeature size and shape distributions, and the effect of changing material strength properties.The use of this modeling approach to probabilistically describe the implications of changes in the microstructure has been demonstrated, thereby allowing the effects of material pedigree to be predictively linked with the structural integrity of end components. The modeling framework has potential applications in airframe design support processes, and as a tool for use in material and product form selection processes.