【摘要】：結(jié)構(gòu)輕量化是航空航天發(fā)展的永恒主題,復(fù)合材料蜂窩夾芯板憑借比強度高、比剛度大、抗失穩(wěn)能力強等優(yōu)越性能,被廣泛應(yīng)用在飛機的主、次承力結(jié)構(gòu)。在制造、裝配和使用過程中,復(fù)合材料蜂窩夾芯板容易受到?jīng)_擊等損傷并需要進行修理。夾芯板結(jié)構(gòu)修理后可能在復(fù)材層合板、粘接膠層和蜂窩芯層出現(xiàn)不同類型的損傷,修理結(jié)構(gòu)的損傷預(yù)測和強度評估已成為復(fù)合材料研究中急需解決的問題。本文針對復(fù)合材料蜂窩夾芯板及其典型的膠接挖補修理結(jié)構(gòu),通過試驗結(jié)合仿真分析的方法,開展?jié)u進損傷分析和強度預(yù)測研究。針對復(fù)合材料層合板膠接挖補修理結(jié)構(gòu),基于復(fù)材層合板CDM模型和膠層CZM模型建立三維漸進損傷分析的有限元仿真模型,基于Hashin準則編寫UMAT程序?qū)崿F(xiàn)復(fù)材纖維和基體的損傷模擬。仿真結(jié)果在膠層、母板和補片的失效模式,以及極限強度方面與試驗結(jié)果相符,為蜂窩夾芯板修理結(jié)構(gòu)仿真分析提供了層合板部分的有效模型。針對典型Nomex?蜂窩芯層進行平壓和剪切力學性能試驗,獲得了等效剛度、強度等材料參數(shù),總結(jié)蜂窩芯層損傷和變形過程分為“彈性-壓潰/剪潰-皺曲-密實/撕裂”4個階段,發(fā)現(xiàn)蜂窩芯層破壞后依然保留部分承載能力,剪切/平壓的剩余強度約占極限強度的40%~45%�；谠囼灲Y(jié)果對Sandwich夾芯板理論中的蜂窩等效模型進行改進,引入蜂窩壓潰剪潰損傷和損傷后剩余強度,建立了有限元損傷分析模型。采用Besant準則作為蜂窩損傷判據(jù),分別編寫USDFLD和UMAT程序?qū)崿F(xiàn)剛度突然退化和漸進退化的兩種損傷演化規(guī)律。通過蜂窩夾芯板三點彎曲的仿真與試驗對比,驗證了模型的可靠性和高效性。最后針對復(fù)合材料蜂窩夾芯板典型膠接挖補修理結(jié)構(gòu),建立三維漸進損傷分析的有限元模型并進行仿真計算。分析發(fā)現(xiàn)夾芯板修理后強度明顯恢復(fù),修理結(jié)構(gòu)在側(cè)壓載荷下未修理區(qū)域發(fā)生局部屈曲,蜂窩和復(fù)材損傷并擴展,導致結(jié)構(gòu)最終失效。模型仿真結(jié)果在失效模式和極限載荷方面與試驗吻合較好。修理參數(shù)仿真分析表明,隨著補片附加層厚度和搭接寬度增加,結(jié)構(gòu)由于附加彎矩提前失穩(wěn),極限載荷降低;挖補角度和補片搭接寬度越小,膠層內(nèi)部剪應(yīng)力越高。建議修理設(shè)計時應(yīng)避免局部剛度改變過大而對結(jié)構(gòu)承載造成的不利影響。
[Abstract]:Structural lightweight is the eternal theme of aerospace development. Composite honeycomb sandwich panels are widely used in aircraft main and secondary load-bearing structures due to their high specific strength, high specific stiffness, strong ability to resist instability, and so on. During manufacture, assembly and use, composite honeycomb sandwich panels are vulnerable to impact damage and need to be repaired. Different types of damage may occur in laminated laminates adhesive layers and honeycomb core layers after repair. The damage prediction and strength evaluation of the repaired structures have become an urgent problem in the study of composite materials. In this paper, progressive damage analysis and strength prediction are carried out for composite honeycomb sandwich panels and their typical glued repair structures. The finite element simulation model of 3D progressive damage analysis is established based on CDM model and CZM model of composite laminates. Based on Hashin criterion, UMAT program is compiled to simulate the damage of composite fiber and matrix. The simulation results are consistent with the test results in the failure mode of rubber layer, motherboard and patch, and the ultimate strength, which provides an effective model for the simulation analysis of honeycomb sandwich panel repair structure. For typical Nomex? The material parameters such as equivalent stiffness, strength and other material parameters are obtained. The damage and deformation process of honeycomb core layer is divided into four stages: elastic-crushing / shear-wrinkling, compacting / tearing. It is found that the honeycomb core layer still retains part of the bearing capacity after failure, and the residual strength of shear / flat compression is about 40% of the ultimate strength. Based on the experimental results, the honeycomb equivalent model in Sandwich sandwich panel theory is improved, and the finite element damage analysis model is established by introducing honeycomb crushing damage and residual strength after damage. The Besant criterion is used as the criterion of honeycomb damage and the USDFLD and UMAT programs are written to realize the two kinds of damage evolution laws of sudden and progressive degradation of stiffness respectively. The reliability and efficiency of the model are verified by comparing the simulation and experiment of honeycomb sandwich panel with three-point bending. Finally, the finite element model of 3D progressive damage analysis is established and simulated for the typical adhesive repair structure of composite honeycomb sandwich panel. It is found that the strength of sandwich panels recovers obviously after repair and local buckling occurs in the unrepaired area of the repaired structure under lateral pressure. The damage and expansion of honeycomb and composite materials lead to the ultimate failure of the structure. The simulation results agree well with the test in failure mode and limit load. The simulation analysis of repair parameters shows that with the increase of the thickness of the additional layer and the width of the lap, the ultimate load of the structure decreases due to the instability of the additional bending moment, and the smaller the mending angle and the width of the lap, the higher the internal shear stress of the adhesive layer. It is suggested that the local stiffness change should be avoided during the repair design and the adverse effect on the bearing capacity of the structure should be avoided.
【學位授予單位】：哈爾濱工業(yè)大學
【學位級別】：碩士
【學位授予年份】：2017
【分類號】：TB33

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