【作者】 钟小丹；

【导师】 陈普会；

【作者基本信息】 南京航空航天大学 ， 飞行器设计， 2013， 博士

【摘要】 复合材料在机身中的应用已成为新一代大型客机的设计发展趋势。本文基于CCAR25部等文件要求，参考NASA等机构的复合材料机身结构设计研究经验，开展了复合材料机身后段结构的综合设计分析工作，并对其中的典型结构细节进行了研究。本文的主要内容包括：(1)在国内首次完成了复合材料机身增压舱段的结构设计，设计过程中除满足适航条例要求外，还考虑了结构的工艺性因素。最终设计方案比金属方案减重约18.5%，可视面积增大了约10.5%。(2)发展了适用于复合材料机身结构的机械连接设计数值分析方法，包括整体-局部单钉模型和整体-局部结构模型分析方法。前者适用于构型及载荷复杂的连接结构，并可用于拉脱载荷失效分析与强度校核；后者可用于精细分析结构规整且载荷形式较为单一连接设计。(3)提出了复合材料机身结构中主要机械连接区域的构型方案，包括机身壁板的环向以及纵向连接等构型方案，并采用本文发展的多钉连接数值分析方法对上述构型方案进行了细节设计分析。(4)研究了复合材料加筋壁板中长桁关键结构细节及离散源损伤容限等问题。通过长桁终止端处采用缩短胶接界面长度的新型混合连接设计的机理研究及参数化分析，指出了该构型在拉伸载荷下提高起裂载荷的设计规律。通过压差载荷下帽型长桁缘条/蒙皮界面破坏的研究，发现内侧圆角处填充物与自由边处的树脂填充对界面初始破坏模式有明显影响；长桁截面参数对界面性能的影响很小。研究了机身壁板离散源损伤问题，分析了缘条/蒙皮间界面性能对蒙皮裂纹扩展和承载能力的影响。(5)建立了基于MPCM方法的常见复合材料工艺工时估算的修正模型，并编制了成本评估系统；通过对构件制造成本中的复杂性因素进行分析，研究了制造经济性对复合材料结构设计的影响规律，探讨了复合材料机身元构件工艺及布局方案选择的经济性原因；结合维修成本估算方法—Liebeck模型，对机身后段的两种设计方案进行了成本评估。(6)通过复合材料机身后段综合设计分析，发现了以下设计规律：机身壁板分缝除工艺条件及成本因素限制外，还应考虑载荷分布的影响；机身壁板的稳定性要求是结构尺寸设计的主要驱动因素；机械连接是结构静强度设计中的关键点；对于窄体客机的机身段应用复合材料时，并不一定具有成本优势；但若在同等舒适性条件下，长寿命(如9万飞行小时)的客机机身应用复合材料时将更具成本优势。(7)参考咨询通报AC20-107B等资料，规划了复合材料机身结构积木式验证试验，设计并参与完成了顶部壁板轴向拉伸试验，试验结果验证了本文的设计分析方法。更多还原

【Abstract】 The application of composite materials in the fuselage has been the tendency for new generationlarge commercial aircraft. Based on CCAR25and other documentations, the integrated design andanalysis of composite fuselage aft section was carried out, and the design experience of researchinstitutions about composite fuselage structures was also referenced in this process，such as NASA.The typical detail structures in the fuselage were also studied. The main features for this paper werestated as follows:(1) Structure design of the composite pressurized fuselage section was completed firstly in China.The requirements of airworthiness regulations were met, and the manufacturability factor was alsoconsidered in the design process. The weight reduction of the final concept was18.5%, and the visiblearea was increased by10.5%compared with the metal concept.(2) The numerical analysis methods were developed for the mechanical joints of the compositefuselage structures, including global-local single fastener model and global-local structure model. Thefirst method was applicable to the joints with complex configurations and loads, and could be used forjoints failure analysis with pull off load. The second method could be used for precise analysis of thejoints with regular structures and a single load.(3) The configuration schemes of the mechanical joint structures were proposed, includinglongitudinal and transverse splices in the composite fuselage, and the detail structures in theseschemes were analyzed by the multiple-bolted joints analysis methods developed in this paper.(4) The key structural details of the stringers and discrete source damage in the composite panelswere studied. The design rules of the novel combined joint design principle at stringer run-outs withreducing bond-line configurations under tensile load, were developed by mechanism and parametricanalysis. The failure analysis of the hat stringer flange/skin interface under the pressure load showedthat the failure models and debond load were relative to the filler in the inner corner and free flangeedge, and the parameters of the stringer cross section had little impact on the interface property. Theimpact of the flange/skin interface on load capacity and crack extension of the skin was studied by thediscrete source damage evolution analysis.(5) Based on MPCM, the time estimation models of the common composite materials processwere amended, and the cost estimation software was developed. The law of influence of manufacturing economy on composite structure design was studied by the complexity factors analysison the manufactory cost, and the economic reasons for selecting process and configuration of thecomposite fuselage structures were investigated. The cost of two design concepts on fuselage sectionwas evaluated with Liebeck model for maintenance cost estimation.(6) The design rules were discovered by the integrated design and analysis of the compositefuselage section, including that: the arc lengths of individual fuselage quadrants were limited not onlyby the processing capability and cost factors, but also by the load distribution of the section; thestability requirement was the primary design driver which determined the size of panels; themechanical joint was the key issue in the static strength design; the application of composite materialsin fuselage structures of narrow body aircraft did not necessarily have a cost advantage comparedwith metal materials; if the aircraft had a longer flight hours life(e.g.90000hrs,), the cost advantage ofcomposite materials used in the fuselage will be greater with the same comfort condition than metalmaterials.(7) The building block tests plane of the composite fuselage structures was drawn up referencingthe document AC20-107B and so on. The crown panel test under the axial load has been designedand completed. and the analysis methods used in the crown panel design process were verified by thetest result.更多还原