机翼机身对接结构三维断裂分析【含图纸、说明书】
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毕 业 设 计（论 文）任 务 书设计（论文）题目：机翼机身对接结构三维断裂分析 学生姓名：张嘉欢 学号：1204201026 专业：机械设计制造及其自动化 所在学院：机电工程学院 指导教师：王芳丽 职称：发任务书日期：年月日 任务书填写要求1毕业设计（论文）任务书由指导教师根据各课题的具体情况填写，经学生所在专业的负责人审查、系（院）领导签字后生效。此任务书应在毕业设计（论文）开始前一周内填好并发给学生。2任务书内容必须用黑墨水笔工整书写，不得涂改或潦草书写；或者按教务处统一设计的电子文档标准格式（可从教务处网页上下载）打印，要求正文小4号宋体，1.5倍行距，禁止打印在其它纸上剪贴。3任务书内填写的内容，必须和学生毕业设计（论文）完成的情况相一致，若有变更，应当经过所在专业及系（院）主管领导审批后方可重新填写。4任务书内有关“学院”、“专业”等名称的填写，应写中文全称，不能写数字代码。学生的“学号”要写全号，不能只写最后2位或1位数字。 5任务书内“主要参考文献”的填写，应按照金陵科技学院本科毕业设计（论文）撰写规范的要求书写。6有关年月日等日期的填写，应当按照国标GB/T 740894数据元和交换格式、信息交换、日期和时间表示法规定的要求，一律用阿拉伯数字书写。如“2002年4月2日”或“2002-04-02”。毕 业 设 计（论 文）任 务 书1本毕业设计（论文）课题应达到的目的： 本毕业设计课题的主要目的是培养学生综合运用所学的基础理论、专业知识和专业基本技能分析和解决实际问题，训练应用ANSYS软件对机翼与机身对接机构进行有限元建模和三维断裂分析的能力，主要包括以下几个方面： 1调查研究、中外文献检索、阅读与翻译的能力； 2综合运用基础理论、专业理论和知识分析解决实际问题的能力； 3查阅和使用专业设计手册的能力； 4应用catia软件和ANSYS软件进行建模和有限元分析的能力； 5撰写设计说明书（论文）的能力。 2本毕业设计（论文）课题任务的内容和要求（包括原始数据、技术要求、工作要求等）： （1） 熟悉并理解机翼机身对接结构与受力特点；（2） 熟悉和掌握三维断裂理论知识；（3） 应用ANSYS软件对机翼机身对接结构进行三维建模和断裂分析。 毕 业 设 计（论 文）任 务 书3对本毕业设计（论文）课题成果的要求包括图表、实物等硬件要求： 1.外文参考资料及译文（附原文）； 2.毕业设计开题报告一份； 3.有限元分析结果分析说明一份； 4主要参考文献： 1 吴相宪，王正为，黄玉堂主编.实用机械设计手册.中国矿业大学出版社，1993. 2 王洪欣，李木，刘秉忠主编.机械设计工程学M.中国矿业大学出版社，2001. 3 唐大放，冯晓宁，杨现卿主编.机械设计工程学M.中国矿业大学出版社，2001. 4 中国纺织大学工程图学教研室等编.画法几何及工程制图.上海科学技术出版社，1997. 5 史美堂主编.金属材料及热处理.上海科学技术出版社，1983. 6 苏翼林主编.材料力学.高等教育出版社，1980. 7 顾崇衔主编.机械制造工艺学.陕西科学技术出版社，1999. 8 詹熙达主编.CATIA V5R20曲面设计教程. 北京：机械工业出版社,2013. 9 詹熙达主编.CATIA V5R20快速入门教程. 北京：机械工业出版社,2011. 10 刘文珽，罗毅，童明波概率损伤容限分析模型研究J航空学报，1993，14（3）：136-139 11 刘文珽等概率断裂力学与概率损伤容限/耐久性M北京航空航天大学出版社，1998. 12 罗毅，黄培彦，刘文珽裂纹扩展寿命安全可靠性分析模型研究J北京航空航天大学学报，2002，28（1）：113-115. 13 杜永恩概率损伤容限分析体系及其关键技术的研究D西安：西北工业大学，2014. 14 K.Y. Lin and A.V. Styuart. Probabilistic approach to damage tolerance design of aircraft composite structures J. Journal of Aircraft, 2007,44(4):1309-1317. 15 Spencer B F，Tang J. Markov Model for fatigue crack growth J. Journal of Engineering Mechanics，1998,114:2134-2157. 毕 业 设 计（论 文）任 务 书5本毕业设计（论文）课题工作进度计划：2015.12.16-2.16.3.9 毕业实习调研，完成开题报告、中英文翻译、论文大纲 2016.3.19-2016.4.25 提交论文草稿，4月中旬中期检查 2016.4.26-2016.5.6 提交论文定稿 2016.5.6-2016.5.13 准备答辩 2016.5.13-2016.5.26 答辩，成绩评定，修改完成最终稿 所在专业审查意见：通过负责人： 2016 年 1 月18 日 毕 业 设 计（论 文）开 题 报 告设计（论文）题目：机翼机身对接结构三维断裂分析 学生姓名：张嘉欢 学号：1204201026 专业：机械设计制造及其自动化 所在学院：机电工程学院 指导教师：王芳丽 职称：年 月日 开题报告填写要求1开题报告（含“文献综述”）作为毕业设计（论文）答辩委员会对学生答辩资格审查的依据材料之一。此报告应在指导教师指导下，由学生在毕业设计（论文）工作前期内完成，经指导教师签署意见及所在专业审查后生效；2开题报告内容必须用黑墨水笔工整书写或按教务处统一设计的电子文档标准格式打印，禁止打印在其它纸上后剪贴，完成后应及时交给指导教师签署意见；3“文献综述”应按论文的框架成文，并直接书写（或打印）在本开题报告第一栏目内，学生写文献综述的参考文献应不少于15篇（不包括辞典、手册）；4有关年月日等日期的填写，应当按照国标GB/T 740894数据元和交换格式、信息交换、日期和时间表示法规定的要求，一律用阿拉伯数字书写。如“2004年4月26日”或“2004-04-26”。5、开题报告（文献综述）字体请按宋体、小四号书写，行间距1.5倍。毕 业 设 计（论文） 开 题 报 告 1结合毕业设计（论文）课题情况，根据所查阅的文献资料，每人撰写不少于1000字左右的文献综述： 机翼机身对接接头设计是飞机结构设计的一个重要环节，其设计的好坏严重关系到飞机的飞行性能和使用安全，本文详细地阐述了飞机结构设计载荷系数的产生和发展，提出在飞机结构设计中用可靠性安全系数替代传统安全系数的观点。介绍分析了机翼机身对接设计思想和接头耳片受力特性，提出了改善机翼机身对接区域传力特性的设计方法。现代超音速战斗机常采用中单翼布置，受机身部位安排的限制，不能有中央翼通过机身，此时机翼通过几个集中接头与机身在侧边相连，机翼弯矩要作用于翼身对接框上，从而机身框结构要加强，结构重量加大。我们知道现代超音速战斗机的机翼很薄，机翼的相对厚度约为4，随着机翼面积的增加，机翼载荷大幅度上升，翼梁承担和传递的载荷越来越大，机翼载荷完全由接头耳片及连接螺栓传递到机身对接框，因此若直接将如此结构高度的翼梁连接到加强框上，机翼机身对接接头耳片、连接螺栓及加强框受力都是比较严重的。由此，现代先进战斗机在总体设计时，结合总体气动布局、机翼机身部位安排的要求采用了多种设计方法来改善对接接头区域的传力特性。目前对接头问题的研究有很多，孙庚茂、丁惠粱等采用钉“超元”模拟复合材料的连接，宋恩鹏等采用梁元和弹簧元分别模拟螺栓连接，张永杰等采用体单元对螺栓连接进行刚度分析。采用螺栓、弹簧元或钉元模拟螺栓连接都是从整体结构承载出发，给出了有限元仿真中螺栓连接的简化方法，但不能解决螺栓的预紧力、孔边接触应力、接头的屈服极限等方面的问题。而大部分接头的出现都会伴随着孔结构的产生孔边的受力由于要考虑摩擦和接触问题，使结构孔边周围呈现出非线性特征，有时为了更大的挖掘接头结构的承载能力，甚至还要考虑结构局部进入塑性区应力分布这就更增加了非线性程度以及求解接头接触问题的复杂性。随着有限元软件的发展，接触问题的数值方法得到了很大的提高，通过不断细化网格可以得到逼近精确解的数值解。本次设计的主要研究工作是探讨传统安全系数的发生和发展，分析传统安全系数的利弊，建立可靠性安全系数观念；对机翼机身对接接头形式和布局以及接头耳片的构型和受力特性进行分析和研究；对原型飞机横梁结构受力特性进行有限元应力分析；根据大改飞机设计要求，优化结构设计参数，给出接头加强型和改进型设计方案；对横梁结构疲劳危险部位进行损伤容限评定。本次设计采用实体建模三维有限元方法，对飞机、翼身对接主承力接头进行传力特性分析，对机身半框模型、机翼主梁一机身横梁组合模型在CATIA和ANSYS下计算结果进行了评价。根据结构设计要求和结构限制条件，优化结构设计参数，给出了飞机机翼机身对接主承力接头的加强型和改进型设计方案。从结构尺寸、重量、疲劳危险部位应力水平、对气动力的影响以及装配工艺性等方面对加强型和改进型设计方案进行了比较分析。疲劳断裂起始于结构细节，对疲劳危险关键部位进行了损伤容限评定，提出了提高结构抗疲劳断裂能力的措施。毕 业 设 计（论文） 开 题 报 告 2本课题要研究或解决的问题和拟采用的研究手段（途径）： 本课题所要研究及解决的问题：本课题是机翼机身对接结构三维断裂分析，需要在给出和查阅到的设计基本资料上完成该装置需要的结构布置，需要研究及解决的问题如下：（1）在了解现有几种三维断裂的原因，需要应用有限元分析软件对断裂原因进行仿真计算；（2）需要利用有限元软件研究分析初始裂纹尺寸的变化，裂纹扩展速率，无损检查，飞行载荷和使用频谱。拟采用的研究手段（途径）：1、文献收集 广泛收集与永磁磁悬浮技术相关的资料； 2、实践与实习 通过实验室内建设基本模型进行测试分析，通过大量的数据来进行设置的计算机控制系统建立。 3、运用CATIA和ANSYS设计软件进行三维分析；4、结合指导老师的指点，分进度，分阶段实施，并对相关问题展开研究。毕 业 设 计（论文） 开 题 报 告 指导教师意见：1对“文献综述”的评语：通过文献综述，该生对机身机翼对接接头三维断裂国内外研究现状有了较清晰的认识，下一步可以通过Ansys软件对对接接头进行有限元建模和三维断裂分析研究。 2对本课题的深度、广度及工作量的意见和对设计（论文）结果的预测：本课题深度和工作量适中，具有一定的工程应用价值，相信通过该生对各种机身机翼对接接头三维断裂的研究，在飞机设计时对对接接头设计时具有一定的参考价值。 3.是否同意开题： 同意 不同意 指导教师： 2016 年 03 月 08 日所在专业审查意见：同意 负责人： 2016 年 03 月 09 日26th ICAF Symposium Montreal, 1-3 June 2011*Challenges in Damage Tolerance Approach forDynamic Loaded Rotorcraft Components FromRisk Assessment to Optimal Inspection PlanningJack Zhao1 and David Adams21 Structural Methods and Prognostics2 Ground TestSikorsky Aircraft CorporationStratford, CT 06516USAjzhaoAbstract. The use of Crack Growth Damage Tolerance as a substantiation methodology for helicopter dynamic components is receiving increased attention as a logical and viable improvement in fatigue reliability and structural integrity. Ithas seen only limited use in helicopters because the addition of difficult periodic inspections was seen as a significant burden to the operator. However the certifying agencies are moving towards the simultaneous use of both Safe-Life and Damage Tolerance methodologies on each component. In order to mitigate the cost issue, a means to optimize the inspection protocol using a risk-informed damage tolerance based fatigue reliability model and maintenance optimization tool is evaluated in this paper. It was desired to maintain the same “6-9s” level ofstructural reliability for Damage Tolerance that is now the standard practice for safe-life substantiations. The newly developed fatigue reliability methodology incorporates the variabilities in initial crack size, crack growth rate, nondestructive inspections, flight loads, and the usage spectrum. The reliability model is further integrated with optimization technique for inspection planning. An example case using the crack propagation test result from a helicopter main rotor spindle is evaluated with the reliability model. The concept of DT risk assessment and optimal inspection planning, impact of NDI detection capability and repair quality on risk reduction, and importance of incorporating CBM logistic requirement are demonstrated. It is concluded that a fatigue reliability model forDamage Tolerance was successfully demonstrated and that it can be used to determine an optimized inspection protocol that reduces the operators inspection burden while providing the required 6-9s level of fatigue reliability.1 IntroductionDamage Tolerance, specifically Crack Growth Damage Tolerance, has been successfully applied in a limited number of helicopter fatigue substantiations for * Oral presentation.928 J. Zhao and D. Adams more than 50 years, although it was originally called “Fail-Safe” methodology. The number of applications is now increasing, driven by an increased emphasis on Damage Tolerance by civil and military certifying agencies. The FAAs Amendment 28 to FAR 29.571 in 1989 provided that Fail-Safety (Damage Tolerance) was an equal-choice option to Safe Life as a substantiation methodology. And a pending new 29.571 will require implementation of both methods on every substantiated component. Damage Tolerance methodology relies on the assumption that the component exhibits some initial damage that subsequently grows progressively over a period of time prior to catastrophic failure. A successful damage tolerance design must be capable of: 1) predicting crack initiation; 2) accurate modelling of subsequent crack growth; and 3) adequate NDI methodology with suitable inspection schedule. The advantage of a Crack Growth Damage Tolerance method over Safe Life is that the cause of an initial crack or damage does not matter since the inspection program will detect the presence of whatever crack occurs before it becomes catastrophic, with a significant safety margin. The disadvantage is the cost of theinspection program in terms of the intrusive down time, man-hours, training, and equipment required. Damage Tolerance will not be accepted as a viable and desirable methodology unless its benefits are perceived to be worth its cost. There is, therefore, an opportunity to employ a reliability approach to determine an optimum inspection methodology one that provides a required level of structural reliability but does not require unnecessary or too-frequent inspections.Conventional Approach to Crack Growth SubstantiationsSikorskys methodology for the substantiation of flight-critical fatigue-loaded components is entirely empirical and was initially developed in the early 1960s for aluminium spar main rotor blades. This substantiation, called “Blade Inspection Method”, or BIM, is still in use today on thousands of rotor blades. It is based on sensing a loss of internal gas pressure in the event of a spar crack, with the inspection interval based on a full-scale fatigue test program that fully characterized the crack growth behaviour under conservative maximum flight loads and severe usage. Sensing of the pressure loss is done by a special visual indicator at the blade root. The inspection interval is essentially a pre-flight visual inspection that was set at minimum 3 to 1 reduction in the test crack growth time from detection to failure. Inspections start at zero time. This method conservative full-scale test determination of crack growth, demonstration of a field inspection method, determination of a failure point, and an inspection interval based on a fraction of the test time is still in use today with a few developments. We now require a static test demonstration for critical crack size, we avoid the inclusion of any blunting effect in metals due to highfatigue test loads, we have employed the method in composites, and we have standard methodology for number of fatigue test specimens and the inspection interval reduction factor. The basic method is accepted by all of our civil and military certifying agencies as illustrated in the figure below from the FAAs AC29-2C MG-11. Damage Tolerance Approach for Dynamic Loaded Rotorcraft Components 929Fig. 1 Potential-to-Functional Failure Curve from NAVAIR 25-403.A reliability determination has not been part of the current crack growth damage tolerance method. Because of the conservative treatments of the flight loads, the usage, and the test-based crack growth characteristics in the substantiation, the current method meets the generic requirement that failure is “extremely remote”, and this criteria has been achieved in 50 years of service. There has been a methodology development, Reference 3, called “Empirical Damage Tolerance”, which allows the determination of an inspection interval for a different load spectrum than was applied in the full-scale test program. This development is also useful in the reliability studies that follow and is described inmore detail later.Reliability-Based Approach to Helicopter Damage ToleranceThe work done to show the reliability of a Damage Tolerant approach for helicopter dynamic component fatigue is not extensive. One early effort did show that a 6-9s level of reliability was achieved for a multiple load path case, Reference 9. However a good starting point for reliability-based approach is Reliability-Centered Maintenance (RCM) as described in NAVAIR 25-403, Reference 5. The figure below illustrates the key points of RCM. This is a much more general methodology, referring to the decline in a functional capability to the point where the functionality is declared failed. The figure is generally known as a P-F curve. The P-F interval is the age interval (in flight hours, cycles, or calendar time) between the Potential Failure (some loss of functionality) becoming detectable (P) to the point of the defined functional failure (F). The inspection interval (I) is a defined fraction of the PF interval. 930 J. Zhao and D. Adams Fig. 2 Potential-to-Functional Failure Curve from NAVAIR 25-403.A reliability-based optimal inspection interval would provide a required predetermined level of structural reliability while minimizing the cost of conducting inspections too frequently. One simplified approach to the reliability is discussed in NAVAIR 25-403, where the Inspection Interval was initially determined by requiring that the projected probability of failure be reduced to less than or equal to the acceptable probability of failure. The interval of on-condition task, denoted as I, can be estimated by:where PF is the Potential-to-Functional failure interval and n denotes the number of inspections during P-F interval. In general, n can be determined by either safety requirements or cost optimization. For flight-critical components, the total riskconsidering the inspections shall not exceed the maximum acceptable risk, Therefore,where acc P is the maximum acceptable level of probability of function failure and is probability of detecting a potential failure in one inspection assuming it exists. The equation above implicitly assumes the failure will always occur in the P-F interval and a constant detectability which is independent to the size of damage. The extreme condition satisfying the risk constraint occurs if the total risk equals to the maximum acceptable level. Accordingly, the number of inspections can be determined by Damage Tolerance Approach for Dynamic Loaded Rotorcraft Components 931 The approach outlined in Eq. 1-3 is based on assumption that a potential failure always exists within the P-F interval and is independent between inspections. As a result, the inspection interval may be too conservative, meaning too-frequent inspections, which does not meet our minimized cost objective. The basic RCM approach does not consider the failure mechanics or the scatter of failure progression. Often, the potential failure mode under consideration exhibits inherent randomness. This is particularly important for the failure modesassociated with progressive damage accumulation such as crack initiation and growth, corrosion, and mechanical wear. To effectively address variability and uncertainty of damage progression and understand their impact on P-F interval, it is highly desirable to incorporate stochastic characterization of failure progression into the RCM process. In this paper, a new approach is proposed to establish a risk-based interval for on- condition tasks by incorporating a baseline probability of failure and a characteristic detectability for inspection capability. Generally, the probability of failure for a component under scheduled inspections can be expressed as the probability of a sequence of events, such as:Where, pF0 is the probability of failure before the first inspection due to excessive damage progression; G i p is the probability that damage will grow to a detectable limit right before the ith inspection (i =1,2, n); ND i p is the conditional probability that inspection will not be able to detect damage at the ith inspection given that damage exists, and F i p is the conditional probability that un-detected damage at the ith inspection will further grow to failure before the next inspection (i+1)th or end of intended service life. Clearly, the probability of failure of these events depends on the probability of damage progress, inspection capability, the timing of inspection, and the number of inspections. Therefore, a more rigorous risk assessment of inspection planning requires comprehensive understanding of the physics of damage initiation, progression and associated randomness, as well as the mathematical model representing inspection capability, and advanced probabilistic methodology capable of performing complicated numerical simulation and assessment. Due to its simplicity for further implementation, the concept of P-F interval and procedure outlined in NAVAIR 00-25-403 serve as a good starting point for establishing a rough estimate of inspection interval. For the purpose of addressing inherent randomness of failure progression and to further facilitate quantitative risk assessment and management for CBM, a more rigorous approach incorporating physics-based damage accumulation model, inspection capability, and advanced probabilistic methodology is needed urgently.932 J. Zhao and D. Adams2 Challenges in a Damage Tolerance ApproachIn damage tolerance approach, structural integrity is ensured through a predictive crack growth model representing the true nature of damage progression, nondestructive inspections to eliminate excessive damaging, and proper repair and maintenance actions. Many factors affects the effectiveness, robustness, an accuracy of the damage tolerance approach, including validation of crack growth model, qualify capability of desirable NDI methods, developing optimal inspection plan, establishing repair limits and criteria for proper maintenanceactions, and setting up rational level of target reliability for risk management. This paper discusses some of the aforementioned technical challenges associated with rotorcraft components and presents a stochastic methodology for predicting rotorcraft component fatigue lifetimes and optimal inspection intervals and assessing underlying risk.Prediction of fatigue crack growth behaviourDamage tolerance approach relies heavily on capability of a fracture mechanics (FM) model to accurately predict potential damage progression initiated at preidentified locations. Several commonly used FM software packages are available for such purpose, including NASGRO, AFGRO, and FASTRAN. They are developed based on linear elastic fracture mechanics and possess a rich library of stress intensity solutions for the commonly encountered structural configuration and geometric profile for anticipated crack growth. From time to time, moreadvanced fracture mechanics may be employed for more complicated crack growth behaviour, structural layups, and loading, if there is the stress intensity solutions do not exist. These advanced fracture mechanics tools, such as BEASY and FRANC-3D, engage boundary element based numerical procedure and simulation. Occasionally, the crack behaviour will also be observed and derived directly from crack growth testing at full component level, such as the empirical damage tolerance approach reported in Reference 3. These approaches represent various levels of modelling and numerical simulation efforts to ensure adequacy of the fracture mechanics model building and accuracy of the predictive capability. For the purpose of qualifying a crack growth model for further DT application, model validation is critical important. There are several ways to achieve the goal. One engages seeded fault testing and the other is to compare the predicted results against the fielded cracking data for further correlation.Uncertainty modeling and quantification for DT approachPrimarily, probabilistic uncertainty analysis and risk assessment involves modeling all of the fundamental quantities entering the problem, and also all uncertainties that arise from lack of knowledge in these quantities, which may affect failure of the component or system. These terms are referred to as basic variables including quantities of structural dimensions and material properties, yield stress and other ultimate response limitations, operating conditions and degradations, environmental and loading factors, etc. The sources of uncertainty in Damage Tolerance Approach for Dynamic Loaded Rotorcraft Components 933 probabilistic analysis can be mainly classified into two categories as aleatory and epistemic uncertainties. Aleatory uncertainties refer to the natural randomness associated with an uncertain quantity, which is inherent in time, in space and measurements. This kind of uncertainty is quantified through the collection and analysis of data to fit to theoretical distributions and, since it is inherent, it cannot be reduced. Epistemic uncertainties reflect a lack of knowledge or information about a quantity, which can be considered in either model or statistical uncertainty-subdivisions. Modal uncertainties arise from simplifications and idealizations that are necessary to model the behavior in a reliability analysis, or from an inadequate understanding of the physical causes and effects. Statistical uncertainties are only due to a shortage of information, and originate from a lack of sufficiently large samples of input data. Statistical uncertainties can be reflectedthrough either parameters of a distribution with a limited set of data or the type of a theoretical distribution to be chosen to fit to data. Since epistemic uncertainty is associated with a lack of knowledge and/or information it follows that it can be reduced through an increase in knowledge by gathering data for a longer period, taking more measurements or carrying out further tests, doing research, and by expert judgment. In order to consider these uncertainties in a structural analysis, appropriate uncertainty models are essential for performing reliability methods to estimate the probability of failure. As one of the key building blocks of a damage tolerance risk assessment and design process, all the sources of uncertainty and their statistical characteristics related to the key design variables must be identified, quantified and further integrated into probabilistic damage tolerance design system. It is well recognized that fatigue initiation and its subsequent crack growth is a random phenomenon. As depicted in Figure 3, various sources of uncertainties contribute to random fatigue and fracture process, including fatigue initiation time, micro-crack initiation and propagation, stress intensity threshold, crack growth rate, usage and loads, and inspection capability for product/in-service inspection.Fig. 3 Uncertainty Identification for Damage Tolerance Approach.It is beyond the scope of this paper to provide comprehensive review of the statistical procedure for modelling of aforementioned sources of variability associated with DT assessment, details of statistical procedure, methodologies, and practices for DT uncertainty identification and modelling can be found inreference 10.934 J. Zhao and D. AdamsProbabilistic risk assessment methodologyProbabilistic methodologies have been widely applied for uncertainty quantification and associated risk assessment. Among the procedures developed for structural reliability assessment and failure probability prediction, a prominent position is held by simulation methods. The Monte Carlo simulation technique, as the basis of all simulation-based techniques, is the most widely applied numerical tool in probabilistic analysis. The convergent rate of the Monte Carlo estimator is appropriately measured by the coefficient of variation of the estimated probability of failure. In general, the basic Monte Carlo technique requires a large sample size to achieve accurate estimate of probability of failure. This becomes a major limitation for the practical application of basic Monte Carlo simulation in structural reliability applications involved in a small probability of failure. To address the challenge, the Importance Sampling technique has been developed and becomes the most prevalent approaches in the context of simulation-based methods for probabilistic analysis. In importance sampling scheme, instead of drawing random samples arbitrarily as the way implemented in a basic Monte Carlo simulation, the majority of the random samples are drawn from the region that contributes the most for the probability of failure. Several approaches can be employed to identify the important region; including 1) MPP obtained through first order reliability methods (FORM) or second order reliability methods(SORM) solution; 2) a priori estimate from pre-sampling; and 3) Markov Chain Monte Carlo simulation. In general, the efficiency of the Importance Sampling technique improves significantly with a large reduction of the variance of estimator, once the appropriate Importance Sampling density function isidentified. In general, DT risk assessment requires generating and repeatable drawings of short-life samples. This requirement dictates the utilization of sampling based methodology in risk assessment and inspection optimization. As the alternative, aMCMC based algorithm has been developed to identified the important region followed by Adaptive Stratified Importance Sampling (ASIS) procedure for the purpose of fast PoF computing and short life sampling (Wu, et.al, 2010).Quantification of NDI capabilityNon-Destructive Inspection (NDI) has been widely used in engineering practice, for both laboratory and field conditions, to ensure adequate structural integrity. Various NDI techniques exist and each of them has its unique capabilities and limitations. Ideally, a perfect NDI should fully detect the presence of a flaw if its size exceeds the detection threshold. In reality, the ideal detection can never be achieved. Due to inherent variability associated with material properties, condition of the structure and its surrounding environment, durability and sensitivity of NDI equipment, field condition to perform inspection, and operator skills, the capability of a NDI is typically expressed by probability of detection (POD) model.Damage Tolerance Approach for Dynamic Loaded Rotorcraft Components 935 Various forms of mathematic representations are available to describe capability of a NDI. Most commonly used POD models are the parametric ones, including Lognormal, Loglogistic, and Gamma distribution. Sometimes, a nonparametric POD model may be used if the detection data cannot be well fitted into the parametric models or lack of sufficient data for a parametric study. Often, a single POD model derived from limited fault detection data which may not be sufficient to represent anticipated fielded conditions. As a result, a lower confidence bound is introduced to address the potential limitation of sample size. In many engineering application, a characteristic crack size represents high detectability with high confidence is used as a indicator of NDI capability. The aforementioned POD modelling choices are illustrated in Figure 4.Fig. 4 Notional Sketch for POD Modelling Options.Once the POD model is fully defined, the probability that growing crack is smaller than a pre-defined characteristic detection threshold can be expressed as where f (x) is the probability density function of growing crack, POD(x) represents the porbability of detection model, and o a is the lower detetion threshold for the POD model. With given P (a) D , the assciated statistical distribution of detections can be estimated bywhere d a is the upper detection limit for the POD model.936 J. Zhao and D. AdamsAccordingly, the probability that growing crack is missed during inspection can be expressed asand statistical distribution of missed detections can be estimated byThese equations are essential for evaluating the statistical distributions of thepopulation of growing crack that are detected or missed during an inspection.在损伤容限的方法动态加载的旋翼机部件风险评估最优检测规划的挑战1.摘要裂纹扩展损伤容限的使用作为一个实体直升机动力部件的方法受到越来越多的关注作为一个逻辑和可行的改善疲劳可靠性和结构的完整性。它在直升机上看到的只有有限的使用，因为增加了困难的周期检查被视为一个重要的负担，操作员。然而认证机构正在走向同时使用两种安全的生活每一个部件的损伤容限方法。为了减轻的成本问题，一种手段，以优化检查协议，使用风险通知基于损伤容限的疲劳可靠性模型及其维修优化在本文中的工具进行评估。这是要求保持同样的“6-9”水平结构可靠性的损伤容限，这是现在的标准做法生命安全验证。新开发的疲劳可靠性方法采用初始裂纹尺寸的变化，裂纹扩展速率，无损检查，飞行载荷和使用频谱。可靠性模型进一步整合优化技术的检验规划。一个例子的情况下，使用裂纹扩展测试结果从直升机的主要用可靠性模型对转子主轴进行了评价。DT风险的概念评估和优化检测规划，NDI检测能力的影响和维修质量的风险降低，并结合煤层气物流的重要性要求证明。它的结论是疲劳可靠性模型成功地证明了损伤容限确定一个优化的检验协议，减少操作者的检查负担的同时提供所需的6-9的疲劳可靠性。引言 损伤容限，特别是裂纹扩展损伤容限，一直是50年以上在有限数量的直升机疲劳验证成功应用，虽然它最初被称为“不安全”的方法。应用程序的数量正在增加，受重视的增加民用和军用认证机构的损坏容限。FAA的修改28到29.571 1989提供故障安全（损伤容限）是一个平等的选择选择安全的生活，作为一种证明方法。和一个等待新的29.571将需要每一个被证实的组件的两种方法的实施。损伤容限方法依赖于假设的组成部分表现出一些最初的损害，随后逐渐增长，在一段时间内灾难性的失败之前的时间。一个成功的损伤容限设计必须能：1）预测裂纹萌生；2）精确建模裂纹扩展；3）足够的NDI方法与合适的检查调度。 裂纹扩展损伤容限法在安全寿命上的优势一个初始裂纹或损坏的原因不重要，因为检查程序将检测到存在的任何裂纹发生之前，它成为灾难性的，具有显着的安全边际。缺点是成本的检查程序中的侵入时间，工时，培训，和设备要求。损坏的宽容将不被接受作为一个可行的和可取的方法，除非它的好处被认为是值得的成本。 因此，有机会采用可靠的方法来确定一种最佳的检验方法，提供了一个要求的水平结构可靠性，但不需要不必要的或过于频繁的检查。传统的方法对裂纹扩展验证西科斯基的方法验证飞行关键疲劳加载组件是完全的经验，最初是在1960年代初开发的铝梁主旋翼叶片。这证明，所谓的“刀片检验方法”，或BIM，是目前仍在使用的转子叶片上的千。它是一种基于梁裂缝事件感知损失的内部气体的压力，以基于全尺寸疲劳试验程序的检验间隔保守最大飞行下裂纹扩展行为的特征负载和严重使用。感测的压力损失是由一个特殊的视觉叶片根部指示器。检查间隔基本上是一个飞行前的视觉在测试裂纹扩展时间内设置至少3到1的检查从检测到故障。检查开始在零时间。示范的现场检查方法，确定故障点，和基于测试时间的一小部分的检验间隔，现在仍在使用有了一些进展。我们现在需要一个静态测试演示的关键裂纹尺寸，我们避免在金属的钝化效果由于高的夹杂物疲劳试验载荷，我们采用的方法在复合材料，我们有疲劳试验样品和检验的标准方法区间折减系数。基本方法是由我们所有的民事和军事认证机构如下图从FAA的交流了mg-11 29-2C。动态加载的旋翼机部件的损伤容限的方法(Damage Size 损伤尺寸, Critical Size 临界尺寸, Threshold of De te ct ability德特能力阈值, Arrested Growth制止增长，No Growth没有增长，Slow Growth缓慢增长，Inspection Interval Basis检验间隔基，time时间)图1 潜在功能失效曲线从NAVAIR 25-403一个可靠的确定还没有一部分的电流裂纹扩展损伤容限法。由于飞行的保守治疗负载，使用和测试为基础的裂纹扩展特性证实，目前的方法满足失败是通用要求“非常遥远”，这一标准已经在50年的服务。有一个方法论的发展，参考 3 ，称为“经验”损伤容限“，它允许一个检查间隔的确定不同的负载频谱比在满量程测试程序中的应用。这发展也很有用的可靠性研究，遵循和描述更详细的细节。基于可靠性的直升机损伤容限方法完成的工作，以显示一个损坏的容错方法的可靠性直升机动力构件疲劳不广泛。一个早期的努力表明一个6-9的可靠性水平是一个多重荷载路径的情况下取得的，参考 9 。然而，一个良好的起点，可靠性为基础的方法是可靠性为中心的维修（RCM）在NAVAIR 25-403描述，参考 5 。下面的图说明了RCM的关键点。这是一个更一般的方法，指的是下降的功能性功能的功能被宣告失败。这个图通常称为P-F曲线。P-F间隔是年龄区间（在飞行时间，周期，或日历时间之间的潜在故障（一些损失功能）成为可检测的（对）的定义的功能的点失败（女）。检查间隔（我）是一个定义的分数的时间间隔。图2 潜在功能失效曲线从NAVAIR 25-403基于可靠性的最优检查间隔将提供所需的预定水平的结构可靠性，同时最大限度地降低成本经常检查的。一个简化的方法的可靠性是在NAVAIR 25-403讨论，在检查间隔最初确定的要求，预测失败的概率减少到少或等于失败的可接受概率。条件的区间任务，记为我，可以估计：在哪里是潜在的功能性故障间隔和氮表示的号码P-F间隔期间检查。在一般情况下，可以由任何一个安全的需求或成本优化。对于飞行关键部件，总风险考虑检查不得超过最大可接受风险，因此，在功能失效概率P是最大可接受水平是检测一检验假设它潜在的失效概存在。上面的方程隐含地假设失败将永远发生在P-F间隔和恒定的探测能力是独立的大小损坏。如果总的情况下，满足风险约束的极端条件发生风险等于最高可接受的水平。因此，数量检查可由该方法在公式1-3中所概述的基础上假设一个潜在的失效总是存在于P-F间隔和检查之间是独立的。作为一个结果，检查间隔可能过于保守，意义太过频繁检查，这不符合我们的最小化成本目标。基本的RCM方法不
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