Material parameter modeling and solution technique using birth–death element for notched metallic panel repaired with bonded composite patch

This paper seeks to outline a novel three-layer model and a new birth–death element solution technique to evaluate static strength of notched metallic panel repaired with bonded composite patch and to optimize material parameters. The higher order 3D, 8-node isotropic solid element and 8-node anisotropic layered solid element with three degrees of freedom per node are respectively implemented to model substrate panel, adhesive layer and composite patch to establish three-layer model of repaired panel. The new solving technique based on birth–death element is developed to allow solution of the stress pattern of repaired panel for identifying failure mode. The new model and its solution are used to model failure mode and residual strength of repaired panel, and the obtained results have a good agreement with the experimental findings. Finally, the influences of material parameter of adhesive layer and composite patch on the residual strength of repaired panel are investigated for optimizing material properties to meet operational and environmental constraints.     

Coupling damage and reliability model of low-cycle fatigue and high energy impact based on the local stress–strain approach

Fatigue induced products generally bear fatigue loads accompanied by impact processes,which reduces their reliable life rapidly. This paper introduces a reliability assessment model based on a local stress–strain approach considering both low-cycle fatigue and high energy impact loads.Two coupling relationships between fatigue and impact are given with effects of an impact process on fatigue damage and effects of fatigue damage on impact performance. The analysis of the former modifies the fatigue parameters and the Manson–Coffin equation for fatigue life based on material theories. On the other hand, the latter proposes the coupling variables and the difference of fracture toughness caused by accumulative fatigue damage. To form an overall reliability model including both fatigue failure and impact failure, a competing risk model is developed. A case study of an actuator cylinder is given to validate this method.     

鸟体撞击结构过程的相似律研究(英文)

With dimensional analysis and similarity theory, the model similarity law of aircraft structures trader bird impact load is investigated. Numerical calculations by means of nonlinear dynamic software ANSYS/LS-DYNA are conducted on the finite element models constructed with different scaling factors. The influence of strain rate on the model similarity law is found to be dependent on the strain rate sensitivity of materials and scale factors. Specifically, materials that are not sensitive to strain rate obey the model similarity law in the bird impact process. The conclusions obtained are supposed to provide a theoretical basis for the experimental work of bird impact on aircraft structure.     

Damage characteristics and constitutive modeling of the 2D C/SiC composite: Part Ⅱ–Material model and numerical implementation

In this work, a macroscopic non-linear constitutive model accounting for damage, inelastic strain and unilateral behavior is proposed for the 2D plain-woven C/Si C composite. A set of scalar damage variables and a new thermodynamic potential expression are introduced in the framework of continuum damage mechanics. In the deduced constitutive equations, the material's progressive damage deactivation behavior during the compression loading is described by a continuous function, and different deactivation rates under uniaxial and biaxial compression loadings are also considered. In damage evolution laws, the coupling effect among the damage modes and impediment effect of compression stress on the development of shear damage in different plane stress states are taken into account. Besides, the general plasticity theory is applied to describing the evolution of inelastic strain in tension and/or shear stress state. The Tsai–Wu failure criterion is adopted for strength analysis. Additionally, the material model is implemented as a user-defined material subroutine(UMAT) and linked to the ABAQUS finite element software, and its performance is demonstrated through several numerical examples.     

A new stress-based multi-scale failure criterion of composites and its validation in open hole tension tests

A new stress-based multi-scale failure criterion is proposed based on a series of off-axis tension tests, and their corresponding fiber failure modes and matrix failure modes are determined at the microscopic level. It is a physical mechanism based, three-dimensional damage analysis criterion which takes into consideration the constituent properties on the macroscopic failure behavior of the composite laminates. A complete set of stress transformation, damage determination and evolution methods are established to realize the application of the multi-scale method in failure analysis. Open-hole tension（OHT） specimens of three material systems（CCF300/5228, CCF300/5428 and T700/5428） are tested according to ASTM standard D5766, and good agreements are found between the experimental results and the numerical predictions. It is found that fiber strength is a key factor influencing the ultimate strength of the laminates, while matrix failure alleviates the stress concentration around the hole. Different matchings of fiber and matrix result in different failure modes as well as ultimate strengths.     

Modeling of a space flexible probe–cone docking system based on the Kane method

Recent developments in micro- and nano-satellites have attracted the interest of the research community worldwide. Many colleges and corporations have launched their satellites in space. Meanwhile, the space flexible probe–cone docking system for micro- and nano-satellites has become an attractive topic. In this paper, a dynamic model of a space flexible probe–cone docking system, in which the flexible beam technology is applied, is built based on the Kane method. The curves of impact force versus time are obtained by the Lagrange model, the Kane model, and the experimental method. The Lagrange model was presented in the reference and verified by both finite element simulation and experiment. The results of the three methods show good agreements on the condition that the beam flexibility and the initial relative velocity change. It is worth mentioning that the introduction of vectorial mechanics and analytical mechanics in the Kane method leads to a large reduction of differential operations and makes the modeling process much easier than that of the Lagrange method. Moreover, the influences of the beam flexibility and the initial relative velocity are discussed. It is concluded that the initial relative velocity of space docking operation should be controlled to a certain value in order to protect the docking system. a 2014 Production and hosting by Elsevier Ltd. on behalf of CSAA BUAA.     

Stress analysis and damage evolution in individual plies of notched composite laminates subjected to in-plane loads

This work aims to investigate local stress distribution, damage evolution and failure of notched composite laminates under in-plane loads. An analytic method containing uniformed boundary equations using a complex variable approach is developed to present layer-by-layer stresses around the notch. The uniformed boundary equations established in series together with conformal mapping functions are capable of dealing with irregular boundary issues around the notch and at infinity. Stress results are employed to evaluate the damage initiation and propagation of notched composites by progressive damage analysis(PDA). A user-defined subroutine is developed in the finite element(FE) model based on coupling theories for mixed failure criteria and damage mechanics to efficiently investigate damage evolution as well as failure modes. Carbon/epoxy laminates with a stacking sequence of [45°/0°/ 60°/90°]sare used to investigate surface strains, in-plane load capacity and microstructure of failure zones to provide analytic and FE methods with strong validation. Good agreement is observed between the analytic method, the FE model and experiments in terms of the stress(strain) distributions, damage evaluation and ultimate strength, and the layerby-layer stress components vary according to a combination effect of fiber orientation and loading type, causing diverse failure modes in individuals.     

Methods to determine stress profile in ALT based on theoretical life models

With the recent products being more reliable,engineers cannot obtain enough failure or degradation information through the design period and even the product lifetime,therefore,accelerated life test(ALT)has become the most popular way to quantify the life characteristics of products.Test design is the most essential topic,such as testing duration,stress profile,data inference,etc.In this paper,a method and procedure based on theoretical life models is proposed to determine the accelerated stress profile.Firstly,the method for theoretical life calculation is put forward based on the main failure mechanism analysis and the theoretical life models.Secondly,the method is provided to determine the accelerated stress profile,including the method to determine the accelerated stress types and the stress range on the basis of the main failure mechanism analysis,the method to determine the acceleration factor and the accelerated stress level based on life quantitative calculation models,and the collaborative analysis method of the accelerated test time while taking the multiple failure mechanisms into consideration.Lastly,the actuator is taken as an example to describe the procedure of the method and the engineering applicability and the validity are verified.     

Uncertainty analysis and design optimization of hybrid rocket motor powered vehicle for suborbital flight

In this paper, we propose an uncertainty analysis and design optimization method and its applications on a hybrid rocket motor(HRM) powered vehicle. The multidisciplinary design model of the rocket system is established and the design uncertainties are quantified. The sensitivity analysis of the uncertainties shows that the uncertainty generated from the error of fuel regression rate model has the most significant effect on the system performances. Then the differences between deterministic design optimization(DDO) and uncertainty-based design optimization(UDO) are discussed. Two newly formed uncertainty analysis methods, including the Kriging-based Monte Carlo simulation(KMCS) and Kriging-based Taylor series approximation(KTSA), are carried out using a global approximation Kriging modeling method. Based on the system design model and the results of design uncertainty analysis, the design optimization of an HRM powered vehicle for suborbital flight is implemented using three design optimization methods: DDO, KMCS and KTSA. The comparisons indicate that the two UDO methods can enhance the design reliability and robustness. The researches and methods proposed in this paper can provide a better way for the general design of HRM powered vehicles.In this paper,we propose an uncertainty analysis and design optimization method and its applications on a hybrid rocket motor(HRM)powered vehicle.The multidisciplinary design model of the rocket system is established and the design uncertainties are quantified.The sensitivity analysis of the uncertainties shows that the uncertainty generated from the error of fuel regression rate model has the most significant effect on the system performances.Then the differences between deterministic design optimization(DDO)and uncertainty-based design optimization(UDO)are discussed.Two newly formed uncertainty analysis methods,including the Kriging-based Monte Carlo simulation(KMCS)and Kriging-based Taylor series approximation(KTSA),are carried out using a global approximation Kriging modeling method.Based on the system design model and the results of design uncertainty analysis,the design optimization of an HRM powered vehicle for suborbital flight is implemented using three design optimization methods:DDO,KMCS and KTSA.The comparisons indicate that the two UDO methods can enhance the design reliability and robustness.The researches and methods proposed in this paper can provide a better way for the general design of HRM powered vehicles.     

A novel method for importance measure analysis in the presence of epistemic and aleatory uncertainties

For structural systems with both epistemic and aleatory uncertainties, research on quantifying the contribution of the epistemic and aleatory uncertainties to the failure probability of the systems is conducted. Based on the method of separating epistemic and aleatory uncertainties in a variable, the core idea of the research is firstly to establish a novel deterministic transition model for auxiliary variables, distribution parameters, random variables, failure probability, then to propose the improved importance sampling(IS) to solve the transition model. Furthermore,the distribution parameters and auxiliary variables are sampled simultaneously and independently;therefore, the inefficient sampling procedure with an ‘‘inner-loop' for epistemic uncertainty and an‘‘outer-loop' for aleatory uncertainty in traditional methods is avoided. Since the proposed method combines the fast convergence of the proper estimates and searches failure samples in the interesting regions with high efficiency, the proposed method is more efficient than traditional methods for the variance-based failure probability sensitivity measures in the presence of epistemic and aleatory uncertainties. Two numerical examples and one engineering example are introduced for demonstrating the efficiency and precision of the proposed method for structural systems with both epistemic and aleatory uncertainties.     

Crashworthiness analysis of aircraft fuselage with sine-wave beam structure

An integrated design concept for crashworthy fuselage using sine-wave beam and strut is proposed and investigated. The finite element model of aircraft fuselage is built first. The structures above cabin floor, occupant and seat are simplified as two rigid blocks. The fuselage frame is redesigned, and the sine-wave beam is arranged under the frame. The impact dynamic performance of the aircraft with bottom sine-wave beam structure is studied and compared with that of conventional type. To obtain better crashworthiness performance, different rigidity of strut is combined with the sine-wave beam bottom structure. Numerical simulation result shows that the proposed sine-wave beam bottom structure could not only dissipate more proportion of impact kinetic energy but also reduce the initial peak acceleration. The structure and rigidity of strut have great influence on the crashworthiness performance. To give a better fuselage structure, both of the strut and bottom structure should be properly integrated and designed.     

大飞机货舱地板下部结构有限元建模与适坠性分析

为了研究大飞机坠撞特性及数值分析方法，选取大飞机货舱地板下部结构为研究对象，建立其有限元模型，实现显式动力学的求解与分析。考察倒置、固支的货舱地板下部结构在200 kg落重以7 m/s垂直冲击下的结构响应、吸能与失效的动态行为，识别落重冲击过程中结构变形与失效模式、冲击响应特性及能量吸收与耗散机理。仿真结果表明，货舱地板下部结构的机身框组件、支撑件组件是主要吸能结构，冲击能量的吸收主要依靠上述结构的塑性变形与失效，紧固件的吸能贡献仅占1%左右。     

客舱地板斜撑杆对民机典型机身段耐撞性能的影响

针对一类具有客舱地板下部斜撑杆的民机典型机身段,研究了斜撑杆对机身段耐撞性能的影响。建立了客舱地板及其下部结构的有限元模型,客舱地板以上的框段结构、乘员质量和座椅被简化为刚性质量块。分析了斜撑杆为开孔/不开孔开剖面结构、斜撑杆剖面尺寸变化以及无斜撑杆和刚性斜撑杆情况下,机身段在坠撞速度为7m/s时的冲击特性。对比分析了各种情况下座椅位置处的加速度-时间历程曲线、机身段的破坏模式和能量吸收情况。结果表明存在合适刚度的斜撑杆,使该类型机身结构具有良好的耐撞毁性能,而增加或减少斜撑杆刚度均有可能产生多次高过载。     

民机机身下部结构耐撞性优化设计

 针对含多设计参数的典型民机机身下部结构耐撞性设计,提出了一种设计方法,该方法以最小化客舱地板的初始加速度峰值与最大化参考压溃状态的结构内能为优化双目标,通过Kriging模型对结构的冲击响应进行预测,采用非支配排序遗传算法II(NSGA-II)对双目标进行优化,进而由Nash-Pareto策略获得最优方案。为了得到最优设计方案,同时研究设计参数对机身结构耐撞性的影响,提出最大化期望提高与最大化预测方差同步加点准则建立代理模型。采用该设计方法,以典型民机机身下部结构设计问题为算例,对客舱地板支撑结构、货舱地板和泡沫构件形状参数进行优化。结果表明,相对原始设计客舱地板的加速度峰值降低约18.3%,次高加速度峰值也得到有效降低,改善了机身结构的耐撞性;Kriging模型预测响应与有限元分析结果误差小于1%,说明了设计方法的有效性。     

民机机身耐撞性设计的波纹板布局

 耐撞性是民机机身结构设计的一项重要要求。为了研究以波纹板为吸能结构的机身结构能量吸收特性和冲击响应特性,针对常规的机身构型,提出了3种货舱地板下部波纹板布局形式,建立了相应的机身段有限元模型,对机身垂直撞击刚性地面的情况进行分析。获得了不同布局形式下的波纹板变形模式和能量吸收情况,以及机身段的破坏模式、能量吸收情况和座椅处的过载-时间历程。3种布局的对比分析表明,在机腹隔框下端和蒙皮之间布置波纹板,可以使机身下部结构的破坏模式稳定,显著降低座椅处的过载峰值,缩短高过载脉宽,有效提高机身结构的耐撞性。     

固体推进剂药柱不确定结构分析及方法比较

针对复合固体推进剂力学性能参数的不确定性对固体火箭发动机药柱结构分析的影响,使用软件ANSYS parametric design language(APDL)建立了受固化降温载荷和压力载荷联合作用下药柱结构的参数化有限元模型.在此基础上,分别应用蒙特卡洛法和响应面法,研究了复合固体推进剂热膨胀系数与初始泊松比的随机分布对药柱结构有限元分析的影响,并对两种方法得到的概率分析结果进行了对比.结果表明:复合推进剂热膨胀系数和初始泊松比微小的变化会对结构分析结果产生较大的影响;药柱结构响应对初始泊松比更为敏感.在药柱结构有限元分析时考虑推进剂力学性能参数的不确定性十分必要.通过对两种不确定结构分析方法的比较发现,响应面法得到的分析结果与蒙特卡洛法得到的分析结果十分接近,且分析效率远高于蒙特卡洛法.      

冲击载荷作用下的锥壁失效仿真及试验验证

根据航空发动机支承锥壁结构受力特点,对风扇叶片飞失冲击载荷作用下的锥壁失效破坏机理进行了研究。利用显式动力学有限元仿真方法,对冲击载荷作用下的锥壁结构动态失效过程进行了瞬态分析。开展了对锥壁的落锤冲击试验,并与分析结果进行了对比验证。试验和分析结果表明:冲击载荷作用下锥壁减薄处破坏为剪切失效破坏。采用的显式动力学有限元仿真方法为准确模拟冲击载荷作用下的锥壁结构失效提供了一种可行的仿真手段,分析获得的峰值加速度与试验结果误差小于5%;利用GISSMO(generalized incremental stress state dependent damage model)可以准确预测锥壁减薄处断裂时间、断裂位置。经过试验验证的分析方法及失效模型可运用到风扇叶片飞失冲击载荷下的锥壁失效设计参数的确定,提高锥壁降载结构设计的精度。      

基于失效概率的涡轮后机匣全局灵敏度分析

涡轮后机匣是航空发动机安全的关键部件,但是其具有工况复杂、不确定性因素多的缺点。为了探究输入随机变量的不确定性对涡轮后机匣结构失效概率的影响,建立参数化有限元模型进行确定性分析。考虑材料性能、几何参数及外部载荷的不确定性,对涡轮后机匣两种典型失效模式：强度失效以及刚度失效建立极限状态函数;通过构造自适应Kriging 代理模型并结合重要抽样方法评估涡轮后机匣结构失效概率,利用基于失效概率的全局灵敏度方法对涡轮后机匣结构可靠度的不确定性来源进行分析,对各输入随机变量重要性进行排序,构建一种涡轮后机匣全局灵敏度分析框架。结果表明：涡轮后机匣在两种失效模式以及系统失效模式下,发动机推力以及线性膨胀系数对结构失效概率影响最为显著,应对其重点考虑;内外机匣长度以及材料弹性模量对涡轮后机匣结构失效概率影响较小,可对其适当忽略。     

Research on the Plane Multiple Cracks Stress Intensity Factors Based on Stochastic Finite Element Method

Taylor stochastic finite element method (SFEM) is applied to analyze the uncertainty of plane multiple cracks stress intensity factors (SIFs) considering the uncertainties of material properties, crack length, and load. The stochastic finite element model of plane multiple cracks are presented. In this model, crack tips are meshed with six-node triangular quarter-point elements; and other area is meshed with six-node triangular elements. The partial derivatives of displacement and stiffness matrix with respect to all the random variables obtained by Taylor SFEM are derived. Meanwhile, the mean and variance expressions of SIF under uncertain factors are also derived. Parallel double-crack illustrative example of using the proposed method is given. The calculation results indicate that the uncertainty of SIF is influenced by the distance of the cracks, the smaller the distance of cracks is, the greater the SIF uncertainty is, and the uncertainties of material elastic modulus, load and crack length markedly affect the uncertainty of SIF; and the Poisson ratio of material has little influence. Among the variables, the elastic modulus has the greatest effect on SIF uncertainty. The next is external load. The crack length has lower effect on SIF uncertainty than both the elastic modulus and external load.     

运输类飞机典型货舱地板下部结构冲击吸能特性

为了研究运输类飞机货舱地板下部结构在冲击载荷作用下的吸能特性，选取三框两段典型货舱地板下部结构试验件开展落重冲击试验，即质量为478.5 kg的落重以3.95 m/s的速度垂直冲击倒置并固定在测力平台上的试验件，分析试验件失效模式及动态响应，同时建立有限元模型进行仿真与试验结果相关性分析及吸能特性研究。结果显示，在此种工况的冲击载荷作用下，中间支撑件发生由32框面向34框面方向的弯曲，并带动机身框发生同向弯曲和扭转，从而导致C型支撑件发生与中间支撑件相反方向的弯曲变形，并最终在机身框与C型支撑件的连接处形成两处塑性铰；紧固件失效以位于中间支撑件附近区域的长桁和剪切角片连接处的22个扁圆头铆钉发生剪切失效为主；试验初始加速度峰值和初始撞击力峰值分别为25.1g和173 kN。仿真与试验获得的结构变形模式吻合较好，仿真获得的最大压缩量与试验结果24.3 mm相差3.7%，仿真获得的压板上初始加速度峰值与试验结果25.1g相差4%。通过仿真分析发现机身框和中间支撑件是主要的吸能部件，吸能贡献分别占总吸能的32.1%和30.4%。