腐蚀环境下复合材料修补件力学特性研究

采用复合材料补片胶接修补含裂纹LY12CZ铝合金板,开展了试验室大气环境和加速预腐蚀环境下复合材料修补件静强度拉伸和疲劳裂纹扩展对比试验研究.结果表明,复合材料补片均能显著提高损伤结构的拉伸强度和疲劳寿命,且短周期的预腐蚀环境对修补件两种力学特性的影响可以忽略不计.同时,基于Paris公式和Rose分析模型,建立了常规环境和预腐蚀环境下疲劳裂纹扩展寿命预测模型,通过与试验结果的对比证明了该模型的工程有效性.     

金属损伤复合材料胶接修补结构力学试验研究

为评估金属损伤复合材料胶接修补结构的剩余强度和剩余寿命,进行了含腐蚀和疲劳损伤LY12CZ航空铝合金板碳/环氧复合材料补片胶接修复结构的力学性能试验,分析了在静拉伸和疲劳载荷作用下修补结构的破坏模式、失效机理。试验研究发现:复合材料胶接修补技术有效改善了腐蚀和疲劳损伤这两类损伤区域的受力状况,恢复了其载荷传递路线,其静拉伸失效形式为金属韧性断裂后胶层脱粘的分步渐进式破坏;与含损伤未修补试样相比,胶接修补大幅度提高了试件的剩余强度和剩余寿命恢复率,修补后铝合金试件承载能力增加了约25%,疲劳寿命增加至修补前的约4倍~5倍。     

复合材料修补结构腐蚀扩展特性实验研究

为了研究腐蚀环境对复合材料修补结构疲劳特性的影响,开展以某型飞机加速环境谱为基础的预腐蚀试验,并采用递进式的试验方法分别完成对未修补未腐蚀、修补但未腐蚀、又修补又腐蚀的试样疲劳裂纹扩展试验。利用试验数据计算了复合材料修补构型、腐蚀环境对裂纹尖端应力强度因子影响的修正系数,并确定了通过试验验证的修补结构疲劳裂纹扩展修正模型。该模型预测腐蚀环境下修补结构的疲劳寿命与试验值一致。试验数据和预测模型可为海军飞机复合材料修补结构的损伤容限设计提供参考。     

7075铝合金扭转微动疲劳行为研究

在基于多轴疲劳试验机上自主设计的扭转微动疲劳装置上,对7075铝合金材料进行了扭转微动疲劳试验,讨论了相同法向载荷下不同扭转切应力对扭转微动疲劳寿命的影响,建立了7075铝合金扭转微动疲劳S-N曲线,并采用SEM,EDS,EPMA等分析方法对扭转微动疲劳的损伤区域进行了分析,建立了扭转微动疲劳混合区接触表面损伤和裂纹萌生及扩展的物理模型,揭示了7075铝合金扭转微动疲劳的损伤机理.试验结果表明:微动作用导致疲劳寿命大大降低;扭转微动疲劳S-N曲线呈"ε"型曲线特征;损伤区靠加载端位置氧化严重,氧化程度随着循环次数增加而增加;微动疲劳的裂纹萌生于次表面,接触区中心两侧主裂纹扩展交叉后垂直于接触表面扩展至试样断裂.     

飞机谱载下铝合金锪窝孔结构腐蚀疲劳研究

通过战斗机谱载下7475—T761铝合金紧固孔试件空气与3．5％NaCl环境中的疲劳试验和断口分析，得到了相应结构疲劳全寿命和孔边角裂纹的裂纹长度-寿命数据．通过三维有限元分析得到了孔边角裂纹的应力强度因子．采用我们新近发展的三维疲劳寿命分析软件，基于材料裂纹扩展基准曲线和一组直通孔实验数据，对锪窝孔结构角裂纹的疲劳和腐蚀疲劳扩展寿命进行了预测．结果表明，提出的基于三维断裂力学的寿命预测方法和软件对复杂结构的全寿命有精确的预测能力，比传统的局部应变法优越，可适用于复杂环境下三维结构的疲劳寿命分析．     

随机疲劳累积损伤理论研究进展

对疲劳载荷、应力、寿命和强度的随机性分别进行了明确分类．然后,针对在疲劳理论发展史上有代表性的、目前在国内外仍被大量引用的（主要用于裂纹起始阶段的）疲劳累积损伤准则,从力学、概率统计等方面进行了较深入系统的分析．提出疲劳损伤累积具有两类不同的概率机制,即损伤在单一个体内具有的随机性以及损伤在同一母体中不同个体之间具有的分散性．最后,重点介绍了二维概率Miper准则以及有关随机疲劳寿命、随机疲劳强度等方面的研究进展．      

裂纹闭合和循环塑性预应变及循环载荷压缩部分对裂纹扩展速率的影响

本文针对三种国产材料 Ｌｙ１１ｃｚ、 Ｌｙ１２ｃｚ 铝合金和 １８ Ｍｎ Ｈ Ｐ钢,通过实验初步考察了循环塑性预应变和循环载荷压缩部分对疲劳裂纹扩展的影响;采用电测法,测定了两种铝合金材料疲劳裂纹扩展的张开应力和有效应力强度因子幅值比 Ｕ。结果表明：（１）材料循环塑性预应变和循环载荷压缩部分,都使疲劳裂纹扩展速率提高;（２）常幅载荷下,在疲劳裂纹稳定扩展阶段,有效应力强度因子幅值比 Ｕ 与应力比 Ｒ 有关,与裂纹长度ａ 无关,并依赖于材料的力学性能。     

含多源损伤的胶接修补结构有限元分析方法

采用有限元法研究含多源损伤结构的胶接修补问题,利用二维三层有限元模型对损伤区进行了数值模拟,并选取典型多源损伤情况中含共线双裂纹铝板结构为算例,详细分析了含多裂纹胶接修补结构中两裂纹相对位置、补片尺寸、铺层和厚度对应力强度因子的影响.结果表明,复合材料胶接修补可明显降低含共线双裂纹母板的应力强度因子;对于确定的裂纹和应力场,应对复合材料补片长度和厚度等参数进行优化设计,以获得最佳的修补效果.     

铝合金切口件在四组变幅载荷下3.5%NaCl溶液中腐蚀疲劳裂纹起始寿命试验研究与估算模型

试验研究了铝合金切口件在4组变幅块谱下、3．5％NaCl溶液中腐蚀疲劳裂纹起始寿命．结果表明，变幅载荷谱中的大超载会显著延长腐蚀疲劳裂纹起始寿命，并且加载顺序具有明显的影响．根据反映大小载荷交互作用的超载腐蚀疲劳裂纹起始寿命曲线和Miner累积损伤定则，建立了变幅载荷下切口件腐蚀疲劳裂纹起始寿命估算模型，应用该模型估算的结果与试验结果相吻合．     

FGHxx粉末高温合金疲劳损伤行为表征、评价与展望

FGHxx系列结构材料是我国近年来成功开发的损伤容限型粉末高温合金,由FGH95到FGH99系列,主要是通过材料的化学成分、制粉工艺及后续的热处理工艺参数等优化措施,达到提高在役温度下的强度、韧性性能,改善疲劳裂纹扩展阻力等目的。相比于材料工艺力学的现有研究成果,该系列材料在固体力学相关的疲劳损伤表征、寿命评价、微结构对疲劳裂纹萌生行为的研究存在显著滞后,尤其在损伤容限相关的关键科学问题以及工程关注的安全评估手段等有待深入探究。本文基于作者们对现有的FGHxx系列粉末高温合金疲劳损伤研究现状和关键科学问题的掌握和理解以及对应的部分实验研究结果,从材料微结构特征、疲劳数据分散特点、宏微观疲劳裂纹扩展行为、SEM原位测试、疲劳寿命预测模型、疲劳损伤安全区（三维K-T关系图）的建立等进行了评述和展望。     

Composite patch reinforcement of cracked aircraft upper longeron: analysis and specimen simulation

The use of composite patches on cracked portions of metallic aircraft structures is an accepted means of improving fatigue life and attaining high structural efficiency. As more and more advanced composite materials are beng developed, the wider use of the repair technology is anticipated even for the reinforcement of primary aircraft structure. The objective of this work is to illustrate how the composite patch repair technology can be successfully applied to restore the structural integrity of cracked components.The Phosphoric Acid Anodize (PAA) surface treatment on aluminum when applied in conjunction with the AVI13/HV998 adhesive were essential for achieving the appropriate patch bonding strength. Such a process was done without immersing the component into the PAA tank; dismantling the component from the aircraft was not necessary. Boron/epoxy and carbon/epoxy patches were applied at room temperature to the 7075-T6511 cracked specimens and tested under fatigue simulating the load spectrum for the upper longeron attached to the access door of the electronic equipment bay. Considerable improvement in the fatigue life was observed after the repair. Equivalent flight test hours were increased from approximately two thousand hours at which the component fractured completely when not repired to twelve thousand hours when the repair was made with only a small amount of crack growth. A six times increase fatigue life is obtained. The laboratory developed technique has been applied to several in-service aircraft which have now been flown for more than 700 h without detection of crack growth.     

复合材料/金属混合结构热应力分布规律

以飞机机身典型部位复合材料与Z型长桁螺栓连接为研究对象,采用Python脚本语言编程,在ABAQUS平台建立了6种紧固长度有限元模型,模拟飞机在高空低温的飞行环境,得到了混合结构的应力、应变分布规律。数值仿真结果表明:铝梁顶部应变结果呈现中间高、两边低的现象,紧固螺栓应变结果则呈现出相反的趋势;末端紧固件承受的剪切载荷最大,而中心的紧固件剪切载荷最小;该温度场结果与已有文献的试验测试结果误差在3℃以内,对复合材料与金属混合结构的设计具有一定意义。     

Strength and stiffness of sandwich beams in bending

This investigation is concerned with the experimental versus analytical correlation of the mechanical properties of sandwich-beam specimens. Such sandwich structures are commonly employed in the aircraft industry. Four-point and three-point load tests were conducted on a large number of sandwich-beam specimens, fabricated by using fiber-glass reinforced plastics (both unidirectional and woven-glass cloth) and DTD 685 aluminum alloy for the facings with aluminum honeycomb core and polyurethane foam cores and the indigenously available Araldite as the bonding medium between the core and the facings.The flexural stiffness of the composite sandwich specimens used in this investigation compared favorably with theoretical predictions. The shear stiffness was found to be about 55 percent and 45 percent of the theoretically predicted values for FRP (fiberglass-reinforced-plastic) cloth and FRP unidirectional laminates with aluminum honeycomb core sandwich, respectively. The failure load as determined by experiments was less than the theoretically predicted safe load. There was a loss of strength as well as a steep decrease in the failure load in the case of low density foam core.It was concluded that FRP facing plates with aluminum honeycomb core sandwich structure may be preferred to similar aluminum-alloy facing sandwich construction if high flexural stiffness and shear stiffness properties are required at less cost and weight. Indigenously available Araldite was quite satisfactory for bonding the core to the facings.This investigation has confirmed the importance of experiments in the field of sandwich structures which can effectively replace other conventional uneconomical structural or machine members which are currently in use.     

Crack Patching: Analysis and Design

Bonded composite patches are frequently used to retard crack growth. This repair procedure is usually referred to as crack patching. The present paper outlines the various methods for the analysis and design of fiber composite patches in thin and thick structures. As illustrative examples the repair of fatigue cracks in the wing skins of Mirage III aircraft, of surface flaws in Macchi landing wheels, and of cracks in a truss is considered.     

Thermal effects on adhesively bonded composite repair of cracked aluminum panels

This study introduces the two-dimensional finite element analysis involving three layer technique to investigate the adhesively bonded composite repair of cracked metallic structure under thermo-mechanical loading. The thermal loading involves, in this study, the temperature drop such as seen during the bonding process. Three patch materials having different stiffnesses and coefficients of thermal expansion are investigated to analyze the thermal effects on the damage tolerance of the crack in the repaired structure and of the debond in the adhesive bondline. For the single sided repair, the patch material having the maximum mismatch in the coefficient of thermal expansion with that of the cracked aluminum plate provides the better damage tolerance capability for both the crack in the panel and the debond in the adhesive. On the other hand, for double sided repair, the patch material having the minimal mismatch in the coefficient of thermal expansion with that of the cracked plate provides the better damage tolerance capability.     

Fatigue crack growth analysis of adhesively repaired panel using perfectly and imperfectly composite patches

Fatigue crack behavior of cracked aluminum panel repaired with the imperfectly bonded composite patch is analyzed. The imperfection is in the form of debond which could result during the bonding of patch or the service life of the repaired structure. Debonds, of various sizes and at different locations with respect to the crack front, are investigated. An analytical procedure, involving two-dimensional finite element method having three layers to model cracked plate, adhesive and composite patch, is used to compute the stress intensity factors of test coupons. From the computed stress intensity factors, the crack growth rates are obtained analytically by assuming that the relationship between the stress intensity factor and the crack growth rate after repair is the same as the fatigue crack growth relationship for cracked panel material. The fatigue crack growth rates obtained experimentally and analytically are in good agreement with each other and they vary linearly with crack length inside the patch. The experimental results are bounded between its analytical counterparts at the mid-plane and free edge surfaces of the cracked panel. The present analytical procedure can, thus, be used to characterize the effects of imperfectly and perfectly bonded composite patch repairs on the durability and damage tolerance of the repaired structure.     

泡沫铝夹层结构抗冲击性能的近场动力学模拟分析

针对某光学舱所采用的泡沫铝夹层防护结构在破片冲击下的抗冲击性能问题,采用Monte-Carlo方法创建了泡沫铝结构的二维细观模型,在常规态型近场动力学理论中引入了Mises屈服准则和线性各向同性强化模型,建立了近场动力学塑性本构的数值计算框架。基于近场动力学计算程序模拟了低速冲击作用下泡沫铝夹层结构的塑性变形以及有机玻璃背板的裂纹扩展形态,分析了泡沫铝芯材孔隙率对该夹层结构抗冲击性能和损伤模式的影响规律。结果表明：泡沫铝夹层结构良好的塑性变形能力是其发挥缓冲与防护作用的主要因素,并且在一定范围内,泡沫铝芯材孔隙率越高,则夹层结构具有更好的抗冲击性能;当泡沫铝孔隙率从0.4提升到0.7时,泡沫铝对冲击物的动能吸收率从90%提高到99%;模拟结果与实验结果具有较好的一致性,验证了模拟结果的准确性和分析结论的有效性。通过数值模拟,预测了有机玻璃背板的裂纹扩展形态,发现提高泡沫铝的孔隙率能获得更好的防护效果。     

Surface hardness increase of 2024 aluminum alloy subjected to cyclic loading

Constant amplitude fatigue tests at R = 0.1, conducted on the aircraft aluminum alloy 2024 T3, have revealed an appreciable surface hardness increase of the alloy at the nano- and meso-scale during fatigue. The observed surface hardness changes could be monitored with confidence by means of nanoindentations. The degree of hardening increases with increasing number of fatigue cycles following exponential relations. With increasing fatigue stress level degree of hardening increases as well. The observed results provide a basis for developing concepts to early detect and also monitor fatigue damage accumulation in aluminum aircraft structures based on measurements of the material’s hardness changes by means of nanoindentations.