【摘要】：在過去的30多年,飛機主要結(jié)構(gòu)上的先進復合材料使用量已顯著增加。隨著先進通用航空運輸試驗計劃的實施,美國國家航空航天局(NASA)和美國聯(lián)邦航空管理局(FAA)在1994年新增了通用和商用航空復合材料的使用。金屬和復合材料在損傷力學和耐久性問題上存在很大的區(qū)別,但復合材料和金屬需要滿足相同要求的結(jié)構(gòu)完整性、安全性和耐用性。雖然復合材料結(jié)構(gòu)的優(yōu)點很多,但由于在大型結(jié)構(gòu)、交互失效機理、溫濕度的敏感性和疲勞分散性等方面缺乏經(jīng)驗,使得其相關試驗論證變得十分困難。本文研究的總體目標是在保證可靠性的同時,考慮經(jīng)濟和試驗持續(xù)時間,采用改進載荷壽命系數(shù)法,求得復合材料的載荷放大系數(shù)(LEF)和試驗持續(xù)時間(N0),并為復合材料的機體結(jié)構(gòu)論證提供指導。具體內(nèi)容包括:(1)概述復合材料飛機結(jié)構(gòu)耐久性試驗,介紹其積木式試驗方法以及分散性分析方法,并且詳細介紹已有復合材料飛機結(jié)構(gòu)耐久性疲勞試驗方法—壽命系數(shù)法、載荷放大系數(shù)法和極限強度法,重點推導改進載荷壽命系數(shù)法,并依此法求得復合材料的LEF和N0,在保證相同可靠性和置信度的前提下,能夠有效地縮短復合材料飛機結(jié)構(gòu)的全尺寸試驗持續(xù)時間。(2)研究鋪層方式為[45/0/-45/90]s的CCF300/BA9916-II復合材料層合板拉-拉載荷的疲勞行為和規(guī)律,使用本文研究的改進載荷壽命系數(shù)法,得到復合材料層合板在B基準上的疲勞壽命和剩余強度的形狀參數(shù),進而求得該復合材料的LEF和N0。并將試驗結(jié)果與已有的復合材料統(tǒng)計結(jié)果相較,驗證改進載荷壽命系數(shù)法的正確性。該方法通過使用改進聯(lián)合威布爾分布分析分散性,簡化了求解復合材料的LEF和N0的過程。(3)研究鋪層方式為[45/-45/0/0/-45/90/0/90/45/0]s的T300/BMP316新型復合材料層合板(無孔和含孔)在拉-拉載荷的疲勞行為及規(guī)律,運用已驗證的改進載荷壽命系數(shù)法求解其LEF和N0,得出部件的試驗件數(shù)和孔徑對復合材料層合板的LEF和N0的影響。
[Abstract]:Over the past 30 years, the use of advanced composite materials in the aircraft's main structures has increased significantly. With the implementation of the Advanced General Air Transport Test Program, NASA (NASA) and Federal Aviation Administration (FAA) added the use of general and commercial aviation composite materials in 1994. There are great differences in damage mechanics and durability between metal and composite materials, but the structural integrity, safety and durability of composites and metals need to meet the same requirements. Although there are many advantages of composite structure, due to the lack of experience in large structure, mutual failure mechanism, sensitivity of temperature and humidity and fatigue dispersion, it is very difficult to test and prove it. The overall goal of this paper is to obtain the load magnification factor (LEF) and the test duration (N0) of the composite material by using the improved load-life coefficient method while ensuring the reliability and taking into account the economic and experimental duration. And provides the guidance for the body structure demonstration of the composite material. The main contents are as follows: (1) the durability test of composite aircraft structure is summarized, and the building block test method and dispersion analysis method are introduced. The life factor method, load magnification factor method and limit strength method are introduced in detail. The improved load life coefficient method is mainly derived. According to this method, the LEF and N0 of composite materials are obtained. With the same degree of reliability and confidence, It can effectively shorten the duration of full-scale test of composite aircraft structures. (2) the fatigue behavior and regularity of tensile and tensile loads of CCF300/BA9916-II composite laminates with [45 / 0 / -45 / 90] s laminates are studied. Using the improved load life coefficient method studied in this paper, the fatigue life and the shape parameters of residual strength of composite laminates on B datum are obtained, and the LEF and N0 of the composite are obtained. The experimental results are compared with the existing statistical results of composite materials to verify the correctness of the improved load life coefficient method. This method uses improved joint Weibull distribution to analyze dispersion. The process of solving the LEF and N0 of composite materials is simplified. (3) A new type of T300/BMP316 composite laminates (without and with holes) is studied in the form of [45 / -45 / 0 / 0 / 45 / 90 / 0 / 90 / 90 / 90 / 45 / 0] s. Fatigue behavior and regularity of tension-pull load, The LEF and N0 are solved by using the improved load life coefficient method, and the effects of the number of test parts and the aperture of the components on the LEF and N0 of composite laminates are obtained.
【學位授予單位】：南昌航空大學
【學位級別】：碩士
【學位授予年份】：2015
【分類號】：V250.2

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1 胡靜偉;余明;萬小朋;;復合材料層板結(jié)構(gòu)疲勞失效數(shù)值仿真研究[J];計算機仿真;2012年04期

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本文編號：2369588