飞机典型薄壁复合材料夹层结构整体屈曲

复合材料学报第35卷 第8期 8月 2018年Acta Materiae Com p ositae Sinica

Vol .35No .8Au g 2018

DOI :10.13801/j .cnki.fhclxb.20180402.006收稿日期:2017-11-23;录用日期:2018-03-27;网络出版时间:2018-04-03 09:39网络出版地址:htt p s ://doi.or g /10.13801/j .cnki.fhclxb.20180402.006基金项目:民用飞机专项科研(MJZ -2014-F -13);黑龙江省青年科学基金(QC2015003)通讯作者:张博明,博士,教授,博士生导师,研究方向为树脂基复合材料力学及工艺Email :zbm@https://www.360docs.net/doc/074984138.html, 引用格式:万玉敏,张发,刘长喜,等.飞机典型薄壁复合材料夹层结构整体屈曲[J ].复合材料学报,2018,35(8):2235-2245.WAN Yumin ,ZHANG Fa ,LIU Chan g xi ,et al.Overall bucklin g of t yp ical thin -wall sandwich com p osites a pp lied on the aircraft [J ].Acta Materiae Com p ositae Sinica ,2018,35(8):2235-2245(in Chinese ).飞机典型薄壁复合材料夹层结构整体屈曲

万玉敏1,张发2,刘长喜3,成强4,孔德拴4,张博明*1

(1.北京航空航天大学材料科学与工程学院,北京100191;2.中国商用飞机有限责任公司北京民用飞机技术研究中心,北京102211;3.黑龙江工程学院机电工程学院,哈尔滨150050;4.中航通飞研究院有限公司,珠海519000)摘 要: 为了研究飞机机身无筋无框复合材料典型薄壁夹层结构在型号上应用的可行性,本文采用解析方法二有限元方法和试验方法对蜂窝夹层复合材料结构的面内压缩和剪切整体屈曲开展系统研究三基于经典层合板理论和工程解析方法推导蜂窝夹层复合材料的压缩和剪切屈曲载荷随试验件尺寸的变化规律三依据某型飞机机身典型结构分别设计压缩和剪切试验件尺寸大小二边界条件和加载方式三利用有限元商用软件ABAQUS 对试验设计建立虚拟试验分析,对比验证解析方法和有限元方法的一致性三最后通过真实试验方法确定解析方法和有限元方法的有效性,并验证典型薄壁夹层结构的承载能力和破坏模式三结果显示,压缩试验结果失效模式与理论预测一致,故3种方法得到的结构整体失稳载荷相近,验证了理论方法的有效性;剪切试验结果发生局部破坏,故试验结果偏低,但有限元方法与解析方法所得结果一致,解析方法相对保守三

关键词: 复合材料;夹层结构;整体屈曲;有限元;解析法

中图分类号: TB330.1 文献标志码: A 文章编号: 1000-3851(2018)08-2235-11Overall bucklin g of t yp ical thin -wall sandwich com p osites a pp lied on the aircraft WAN Yumin 1,ZHANG Fa 2,LIU Chan g xi 3,CHENG Qian g 4,KONG Deshuan 4,ZHANG Bomin g *1(1.School of Materials Science and En g ineerin g ,Beihan g Universit y ,Bei j in g 100191,China ;2.Bei j in g Aeronautical Science &Technolo gy Research Institute ,COMAC ,Bei j in g 102211,China ;3.Colle g e of Mechanical and Electrical En g ineerin g ,Heilon gj ian g Institute of Technolo gy ,Harbin 150050,China ;4.China Aviation Industr y General Aircraft Co.Ltd.

,Zhuhai 519000,China )Abstract : In order to stud y the feasibilit y of a pp l y in g t yp ical thin -wall sandwich com p osite without ribs and frames reinforcement on the aircraft fusela g e ,overall bucklin g p erformance of the hone y comb sandwich com p osite under the in -p lane com p ression and shear load was studied usin g anal y tical method ,finite element method (FEM )and ex p eri -ment.Based on the classical laminated p late theor y and the en g ineerin g anal y tical method ,the variation of the buck -lin g load of hone y comb sandwich com p osites with the sam p le size was g iven and com p ared with the finite element results.Two t yp ical sizes of sandwich structure were desi g ned ,and boundar y conditions and loadin g methods were defined based on a certain t yp e of t yp ical aircraft fusela g e structure.Finall y ,the bucklin g loads obtained b y anal y tic solution and FEM were com p ared with the ex p erimental results ,which verified the bearin g ca p acit y and failure mode of the t yp ical thin -wall sandwich structure.The results show that the test com p ression failure mode is consistent with the theoretical p rediction ,so the overall bucklin g load calculated throu g h the three methods is almost e q ual to each other.Because local failures occur in the shear test ,so the ex p erimental result is lower than FEM result and anal y tic result.However ,the FEM results and the anal y tical results are consistent ,which could verif y their effec -

tiveness.And it s found that the anal y tical method is relativel y conservative.Ke y words : com p osites ;sandwich structure ;overall bucklin g ;finite element method (FEM )

;anal y tical solution 万方数据