飞机结构的腐蚀损伤及其对寿命的影响
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摘要
腐蚀损伤严重地影响着飞机结构的安全性。本论文对飞机铝合金结构的腐蚀损伤及腐蚀对结构寿命的影响从试验和理论角度进行了深入的研究。以实验为基础，对腐蚀损伤形貌进行归类，提出腐蚀损伤的评价指标，系统研究了飞机铝合金结构中广泛存在的坑蚀、剥蚀及腐蚀疲劳损伤，并建立了相关的数学评价模型。给出了一整套的飞机结构腐蚀损伤分析方法及寿命评估技术，力图为腐蚀环境下服役的老龄飞机结构定寿、延寿工作提供尽可能多的技术支持。
从实验角度上来说，首先对LY12CZ铝合金在实验室环境下进行了预腐蚀—疲劳加载试验。在进行预腐蚀试验时，分别设置了三组不同的腐蚀时间和腐蚀温度。对试验件的腐蚀损伤形貌及疲劳断口利用光学显微镜和扫描电镜进行了观测及拍照。试验结果表明：随着腐蚀温度的升高及腐蚀时间的延长，腐蚀坑深度和腐蚀坑表面积都逐渐加大，相邻腐蚀坑相互交错形成更大更深的腐蚀坑，而且伴随着试件表面剥蚀损伤的产生。经过预腐蚀后的铝合金试件在进行疲劳加载时，裂纹均起始于腐蚀坑，大多数试件的断裂由一条主裂纹扩展而导致的，只有部分试件观测到多条裂纹的扩展。在试验件的中央横截面部位，应力集中程度是最严重的，但是在中央部位断裂的试验件并不是很多，而在靠近腐蚀区域和未腐蚀区域交界的地方发生断裂的试验件较多。腐蚀缩短了裂纹的起始阶段，使试验件疲劳寿命降低1.25到2.38个因子。
其次，对LY12CZ含紧固孔结构件腐蚀疲劳裂纹扩展规律进行了试验研究。将含中心通孔的板状试验件在环境条件分别为空气和3.5％NaCl腐蚀溶液中进行疲劳加载。试验表明：腐蚀液中的疲劳裂纹扩展速率大于空气中的扩展速率。频率对腐蚀疲劳裂纹扩展速率有一定的影响，低的加载频率导致较高的裂纹扩展速率。3.5％NaCl腐蚀溶液中的疲劳裂纹扩展速率大于空气中，但是相差较小，这可能是由于腐蚀疲劳过程进行时间太短，从而造成在试验过程中，腐蚀作用对疲劳过程的贡献不大。
从理论角度上来说
(1)提出遭受腐蚀损伤铝合金结构的疲劳性能评价指标是：腐蚀损伤平均深度和表面腐蚀损伤度。讨论了腐蚀损伤形态对疲劳寿命的影响，从腐蚀损伤平均深度和腐蚀损伤度两个方面可以来确定受腐蚀损伤飞机构件的危险部位。
(2)提出利用AB参数可以方便地对蚀坑形貌进行归类。通过对腐蚀损伤形貌的试验观测，引入AB参数来描述腐蚀损伤形貌的演化。随着腐蚀作用时间(温度)的增长，AB值逐渐变大并趋向于0.7854(蚀坑呈圆形或椭圆形)，不规则的蚀坑逐渐形成发展成面积，深度更大的圆形(椭圆)蚀坑。
(3)建立了腐蚀损伤演化的数学模型。对表面腐蚀损伤度数据进行了统计分析处理，指出不同腐蚀温度和不同腐蚀时间下试验件损伤度的概率分布均符合Logistic分布。利用此概率模型可以方便地预测不同服役环境下腐蚀损伤。综合利用电化学、概率统计以及随机过程方法建立了铝合金结构腐蚀损伤演化的数学模型并利用试验数据进行验证。
(4)提出遭受坑蚀及剥蚀损伤结构件的寿命及剩余强度预测方法。以试验测量数据及腐蚀损伤形貌为基础，应用ANSYS分析软件，采用单元生死方法，建立遭受腐蚀损伤试验件的三维有限元模型，分析蚀坑的局部应力情况，考察蚀坑对试验件应力集中的影响程度。并将腐蚀坑简化为半椭圆形三维表面裂纹形式，采用AFGROW软件进行寿命估算及剩余强度研究，并利用试验数据进行了验证，预测精度良好。
(5)提出铝合金结构件腐蚀疲劳裂纹虚拟扩展技术。首次对紧固孔结构件腐蚀疲劳裂纹的虚拟扩展技术进行了研究，提供了随机过程方法和AFGROW模拟两种模拟技术。给定循环数下的裂纹长度值符合对数正态分布，这一结论无论在腐蚀或者非腐蚀环境下，取正态随机过程或者对数正态随机过程时均是成立的。
(6)提出沿海使用飞机结构腐蚀一寿命管理方法。利用我国在役飞机机体腐蚀损伤调研所获取的实际信息，计算了实际使用中腐蚀损伤可允许程度的限定，对坑蚀和剥蚀两种腐蚀形式进行了详细的讨论分析，获得了易于工程应用的简洁的图表。最后提出了腐蚀环境下服役的飞机铝合金结构全寿命预测的模型框架。
Corrosion, in its various forms, was a serious issue to be considered in its effects on aircraft structural safety. The following dissertation described the experimental and theoretical study on the corrosion damage and fatigue life of aluminum alloy structure. Based on the experiments, morphological classification on corrosion damage of aluminum alloy was studied. The metric for evaluating the corrosion damage grade was proposed. The pitting corrosion, exfoliation corrosion and corrosion fatigue which were the most usually occurred on aircraft structures were studied systemically. The primary mathematical relationship of the three kinds of damages and the degradation of fatigue life were built. The methods of analyzing the corrosion damage of aircraft structure and the technique of evaluating the life were developed. The dissertation will provide some technical support to life assessment and life extension of aging aircraft structures.
Two experiments were conducted in this dissertation. Firstly, LY12CZ aluminum alloy specimens were pre-corroded under the conditions of three different test temperatures and exposure durations. After corrosion exposure, fatigue tests were performed. Scanning electron microscopy and optical microscope analyses on corrosion damage were carried out. Experimental results showed that, with the lengthening in corrosion exposure duration and the corrosion temperature going up, the dimension of corrosion pits became larger and little corrosion pits were congregated together, and the corrosion damage expanded along the specimen surface and cross section directions. The near corrosion pits congregated together and the exfoliation corrosion were observed. It can be found that all of the cracks nucleated from corrosion pits. In general, there were multiple fatigue cracks for almost every specimen, but most fatigue failures were governed by a dominant pit. In the middle of the specimens, the stress concentration was the most serious, but the fatigue crack initiating seldom occurred at this place and fracture usually occurred at the borderline of corrosion and non-corrosion. The artificially produced corrosion damage considerably reduced the fatigue life of laboratory specimens because of shortening the crack nucleation stage and the corrosion damage decreased the fatigue lives by a factor of about 1.25 to 2.38.
Secondly, The fatigue crack propagation behavior of LY12CZ aluminum alloy specimen with central hole of fastener has been investigated in air and 3.5%NaCl solution. Experimental results showed that, the corrosion fatigue crack growth rate decreased with the increasing of the loading frequencies, and in corrosive environment, the crack growth rate was slightly, larger than in air. In 3.5 %NaCl, the crack growth rate was slightly larger than in air and the reason maybe rest on the duration of corrosion fatigue test was too short, and this led to the role of corrosion process was not significant.
Based on the experiments, some mathematical models were proposed as follows,
(1) The image analyses indicated that the two key parameters (the depth of corrosion pit and the surface corrosion damage ratio) were the metrics for evaluating the corrosion damage grade. The effect of corrosion damage on fatigue lives was discussed. Using the metrics can fix on the critical place of the aircraft structure subjected to corrosion damage.
(2) The AB (Area-Box) parameter was proposed to conveniently classify the configuration of corrosion pits. Based on the experimental observation of corrosion pits configuration, the proposed AB parameter can describe the evolution rhythm of corrosion pit. With the lengthening in corrosion exposure duration and rise of corrosion temperature, AB value tend to 0.7854 (the corrosion pit presented rotundity or ellipse) and the irregular corrosion pits gradually developed to lager circular or elliptical corrosion pits.
(3) The mathematical model of corrosion damage evolution was proposed. Statistical analysis had been performed to draw the conclusion that the Logistic distribution was acceptable for the data sets of the corrosion damage ratio. Using the probability model of corrosion damage evolution, we can forecast corrosion damage in varied service environments. Combing the electrochemistry, probabilistic analysis and stochastic processes method, the mathematic model of corrosion damage evolution was built and the predicted results were in good agreement with experimental results.
(4) The prediction method of residual strength and lives of corroded structure was presented. Based on the experimental observation, using the element life-death technique in ANSYS software, the 3D finite element model of corroded structure was built and the stress distribution was analyzed. A simple crack growth calculation method using AFGROW software successfully predicted the fatigue lives of the artificially corroded specimens using the depth and width of the corrosion pits as the initial crack size. The simulation results using AFGROW were in good agreement with the experimental data.
(5) The virtual corrosion fatigue crack propagation tests on aluminum structure were developed. The virtual corrosion fatigue crack propagation of fastening hole structure were investigated. The stochastic processes method and AFGROW simulation method were presented. The probabilistic model of crack length for a given number of cycles was well approximated by a lognormal distribution. This was true for either normal or lognormal stochastic processes and in corrosive or non-corrosive environment.
(6) The corrosion-life management method of aircraft structure served in inshore circumstance was proposed. Based on the practical corrosion damage information derived from the investigation on the served aircraft structures, the corrosion damage tolerance of aircraft structure was investigated. The pitting corrosion and exfoliation corrosion were studied particularly and the terse engineering schematics were provided. At last, the holistic life model of aging aircraft made of aluminum alloys served in corrosive environment was proposed schematically.
从实验角度上来说，首先对LY12CZ铝合金在实验室环境下进行了预腐蚀—疲劳加载试验。在进行预腐蚀试验时，分别设置了三组不同的腐蚀时间和腐蚀温度。对试验件的腐蚀损伤形貌及疲劳断口利用光学显微镜和扫描电镜进行了观测及拍照。试验结果表明：随着腐蚀温度的升高及腐蚀时间的延长，腐蚀坑深度和腐蚀坑表面积都逐渐加大，相邻腐蚀坑相互交错形成更大更深的腐蚀坑，而且伴随着试件表面剥蚀损伤的产生。经过预腐蚀后的铝合金试件在进行疲劳加载时，裂纹均起始于腐蚀坑，大多数试件的断裂由一条主裂纹扩展而导致的，只有部分试件观测到多条裂纹的扩展。在试验件的中央横截面部位，应力集中程度是最严重的，但是在中央部位断裂的试验件并不是很多，而在靠近腐蚀区域和未腐蚀区域交界的地方发生断裂的试验件较多。腐蚀缩短了裂纹的起始阶段，使试验件疲劳寿命降低1.25到2.38个因子。
其次，对LY12CZ含紧固孔结构件腐蚀疲劳裂纹扩展规律进行了试验研究。将含中心通孔的板状试验件在环境条件分别为空气和3.5％NaCl腐蚀溶液中进行疲劳加载。试验表明：腐蚀液中的疲劳裂纹扩展速率大于空气中的扩展速率。频率对腐蚀疲劳裂纹扩展速率有一定的影响，低的加载频率导致较高的裂纹扩展速率。3.5％NaCl腐蚀溶液中的疲劳裂纹扩展速率大于空气中，但是相差较小，这可能是由于腐蚀疲劳过程进行时间太短，从而造成在试验过程中，腐蚀作用对疲劳过程的贡献不大。
从理论角度上来说
(1)提出遭受腐蚀损伤铝合金结构的疲劳性能评价指标是：腐蚀损伤平均深度和表面腐蚀损伤度。讨论了腐蚀损伤形态对疲劳寿命的影响，从腐蚀损伤平均深度和腐蚀损伤度两个方面可以来确定受腐蚀损伤飞机构件的危险部位。
(2)提出利用AB参数可以方便地对蚀坑形貌进行归类。通过对腐蚀损伤形貌的试验观测，引入AB参数来描述腐蚀损伤形貌的演化。随着腐蚀作用时间(温度)的增长，AB值逐渐变大并趋向于0.7854(蚀坑呈圆形或椭圆形)，不规则的蚀坑逐渐形成发展成面积，深度更大的圆形(椭圆)蚀坑。
(3)建立了腐蚀损伤演化的数学模型。对表面腐蚀损伤度数据进行了统计分析处理，指出不同腐蚀温度和不同腐蚀时间下试验件损伤度的概率分布均符合Logistic分布。利用此概率模型可以方便地预测不同服役环境下腐蚀损伤。综合利用电化学、概率统计以及随机过程方法建立了铝合金结构腐蚀损伤演化的数学模型并利用试验数据进行验证。
(4)提出遭受坑蚀及剥蚀损伤结构件的寿命及剩余强度预测方法。以试验测量数据及腐蚀损伤形貌为基础，应用ANSYS分析软件，采用单元生死方法，建立遭受腐蚀损伤试验件的三维有限元模型，分析蚀坑的局部应力情况，考察蚀坑对试验件应力集中的影响程度。并将腐蚀坑简化为半椭圆形三维表面裂纹形式，采用AFGROW软件进行寿命估算及剩余强度研究，并利用试验数据进行了验证，预测精度良好。
(5)提出铝合金结构件腐蚀疲劳裂纹虚拟扩展技术。首次对紧固孔结构件腐蚀疲劳裂纹的虚拟扩展技术进行了研究，提供了随机过程方法和AFGROW模拟两种模拟技术。给定循环数下的裂纹长度值符合对数正态分布，这一结论无论在腐蚀或者非腐蚀环境下，取正态随机过程或者对数正态随机过程时均是成立的。
(6)提出沿海使用飞机结构腐蚀一寿命管理方法。利用我国在役飞机机体腐蚀损伤调研所获取的实际信息，计算了实际使用中腐蚀损伤可允许程度的限定，对坑蚀和剥蚀两种腐蚀形式进行了详细的讨论分析，获得了易于工程应用的简洁的图表。最后提出了腐蚀环境下服役的飞机铝合金结构全寿命预测的模型框架。
Corrosion, in its various forms, was a serious issue to be considered in its effects on aircraft structural safety. The following dissertation described the experimental and theoretical study on the corrosion damage and fatigue life of aluminum alloy structure. Based on the experiments, morphological classification on corrosion damage of aluminum alloy was studied. The metric for evaluating the corrosion damage grade was proposed. The pitting corrosion, exfoliation corrosion and corrosion fatigue which were the most usually occurred on aircraft structures were studied systemically. The primary mathematical relationship of the three kinds of damages and the degradation of fatigue life were built. The methods of analyzing the corrosion damage of aircraft structure and the technique of evaluating the life were developed. The dissertation will provide some technical support to life assessment and life extension of aging aircraft structures.
Two experiments were conducted in this dissertation. Firstly, LY12CZ aluminum alloy specimens were pre-corroded under the conditions of three different test temperatures and exposure durations. After corrosion exposure, fatigue tests were performed. Scanning electron microscopy and optical microscope analyses on corrosion damage were carried out. Experimental results showed that, with the lengthening in corrosion exposure duration and the corrosion temperature going up, the dimension of corrosion pits became larger and little corrosion pits were congregated together, and the corrosion damage expanded along the specimen surface and cross section directions. The near corrosion pits congregated together and the exfoliation corrosion were observed. It can be found that all of the cracks nucleated from corrosion pits. In general, there were multiple fatigue cracks for almost every specimen, but most fatigue failures were governed by a dominant pit. In the middle of the specimens, the stress concentration was the most serious, but the fatigue crack initiating seldom occurred at this place and fracture usually occurred at the borderline of corrosion and non-corrosion. The artificially produced corrosion damage considerably reduced the fatigue life of laboratory specimens because of shortening the crack nucleation stage and the corrosion damage decreased the fatigue lives by a factor of about 1.25 to 2.38.
Secondly, The fatigue crack propagation behavior of LY12CZ aluminum alloy specimen with central hole of fastener has been investigated in air and 3.5%NaCl solution. Experimental results showed that, the corrosion fatigue crack growth rate decreased with the increasing of the loading frequencies, and in corrosive environment, the crack growth rate was slightly, larger than in air. In 3.5 %NaCl, the crack growth rate was slightly larger than in air and the reason maybe rest on the duration of corrosion fatigue test was too short, and this led to the role of corrosion process was not significant.
Based on the experiments, some mathematical models were proposed as follows,
(1) The image analyses indicated that the two key parameters (the depth of corrosion pit and the surface corrosion damage ratio) were the metrics for evaluating the corrosion damage grade. The effect of corrosion damage on fatigue lives was discussed. Using the metrics can fix on the critical place of the aircraft structure subjected to corrosion damage.
(2) The AB (Area-Box) parameter was proposed to conveniently classify the configuration of corrosion pits. Based on the experimental observation of corrosion pits configuration, the proposed AB parameter can describe the evolution rhythm of corrosion pit. With the lengthening in corrosion exposure duration and rise of corrosion temperature, AB value tend to 0.7854 (the corrosion pit presented rotundity or ellipse) and the irregular corrosion pits gradually developed to lager circular or elliptical corrosion pits.
(3) The mathematical model of corrosion damage evolution was proposed. Statistical analysis had been performed to draw the conclusion that the Logistic distribution was acceptable for the data sets of the corrosion damage ratio. Using the probability model of corrosion damage evolution, we can forecast corrosion damage in varied service environments. Combing the electrochemistry, probabilistic analysis and stochastic processes method, the mathematic model of corrosion damage evolution was built and the predicted results were in good agreement with experimental results.
(4) The prediction method of residual strength and lives of corroded structure was presented. Based on the experimental observation, using the element life-death technique in ANSYS software, the 3D finite element model of corroded structure was built and the stress distribution was analyzed. A simple crack growth calculation method using AFGROW software successfully predicted the fatigue lives of the artificially corroded specimens using the depth and width of the corrosion pits as the initial crack size. The simulation results using AFGROW were in good agreement with the experimental data.
(5) The virtual corrosion fatigue crack propagation tests on aluminum structure were developed. The virtual corrosion fatigue crack propagation of fastening hole structure were investigated. The stochastic processes method and AFGROW simulation method were presented. The probabilistic model of crack length for a given number of cycles was well approximated by a lognormal distribution. This was true for either normal or lognormal stochastic processes and in corrosive or non-corrosive environment.
(6) The corrosion-life management method of aircraft structure served in inshore circumstance was proposed. Based on the practical corrosion damage information derived from the investigation on the served aircraft structures, the corrosion damage tolerance of aircraft structure was investigated. The pitting corrosion and exfoliation corrosion were studied particularly and the terse engineering schematics were provided. At last, the holistic life model of aging aircraft made of aluminum alloys served in corrosive environment was proposed schematically.
引文
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