Structure Design Of Composite Pressurized Fuselage Section In Commercial Aircraft

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.