復(fù)合材料疲勞試驗(yàn)的改進(jìn)載荷壽命系數(shù)法研究
發(fā)布時(shí)間：2018-12-09 15:08
【摘要】：在過去的30多年,飛機(jī)主要結(jié)構(gòu)上的先進(jìn)復(fù)合材料使用量已顯著增加。隨著先進(jìn)通用航空運(yùn)輸試驗(yàn)計(jì)劃的實(shí)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(fù)合材料的使用。金屬和復(fù)合材料在損傷力學(xué)和耐久性問題上存在很大的區(qū)別,但復(fù)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(fù)合材料結(jié)構(gòu)的優(yōu)點(diǎn)很多,但由于在大型結(jié)構(gòu)、交互失效機(jī)理、溫濕度的敏感性和疲勞分散性等方面缺乏經(jīng)驗(yàn),使得其相關(guān)試驗(yàn)論證變得十分困難。本文研究的總體目標(biāo)是在保證可靠性的同時(shí),考慮經(jīng)濟(jì)和試驗(yàn)持續(xù)時(shí)間,采用改進(jìn)載荷壽命系數(shù)法,求得復(fù)合材料的載荷放大系數(shù)(LEF)和試驗(yàn)持續(xù)時(shí)間(N0),并為復(fù)合材料的機(jī)體結(jié)構(gòu)論證提供指導(dǎo)。具體內(nèi)容包括:(1)概述復(fù)合材料飛機(jī)結(jié)構(gòu)耐久性試驗(yàn),介紹其積木式試驗(yàn)方法以及分散性分析方法,并且詳細(xì)介紹已有復(fù)合材料飛機(jī)結(jié)構(gòu)耐久性疲勞試驗(yàn)方法—壽命系數(shù)法、載荷放大系數(shù)法和極限強(qiáng)度法,重點(diǎn)推導(dǎo)改進(jìn)載荷壽命系數(shù)法,并依此法求得復(fù)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(fù)合材料飛機(jī)結(jié)構(gòu)的全尺寸試驗(yàn)持續(xù)時(shí)間。(2)研究鋪層方式為[45/0/-45/90]s的CCF300/BA9916-II復(fù)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(jìn)載荷壽命系數(shù)法,得到復(fù)合材料層合板在B基準(zhǔn)上的疲勞壽命和剩余強(qiáng)度的形狀參數(shù),進(jìn)而求得該復(fù)合材料的LEF和N0。并將試驗(yàn)結(jié)果與已有的復(fù)合材料統(tǒng)計(jì)結(jié)果相較,驗(yàn)證改進(jìn)載荷壽命系數(shù)法的正確性。該方法通過使用改進(jìn)聯(lián)合威布爾分布分析分散性,簡(jiǎn)化了求解復(fù)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(fù)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(yùn)用已驗(yàn)證的改進(jìn)載荷壽命系數(shù)法求解其LEF和N0,得出部件的試驗(yàn)件數(shù)和孔徑對(duì)復(fù)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(xué)位授予單位】：南昌航空大學(xué)
【學(xué)位級(jí)別】：碩士
【學(xué)位授予年份】：2015
【分類號(hào)】：V250.2
本文編號(hào)：2369588
[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(xué)位授予單位】：南昌航空大學(xué)
【學(xué)位級(jí)別】：碩士
【學(xué)位授予年份】：2015
【分類號(hào)】：V250.2
【參考文獻(xiàn)】
相關(guān)期刊論文 前1條
1 胡靜偉;余明;萬小朋;;復(fù)合材料層板結(jié)構(gòu)疲勞失效數(shù)值仿真研究[J];計(jì)算機(jī)仿真;2012年04期
，本文編號(hào)：2369588
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