TC4 titanium alloy bolted structures are extensively utilized in aerospace engineering, particularly within the heat-affected zones of aircraft engines. However, current studies have predominantly focused on fatigue fracture of titanium alloys at temperatures exceeding 400 °C, leaving a gap in accurate fatigue life prediction for TC4 bolted structures subjected to moderate elevated temperatures up to 400 °C. To address this limitation, this study proposes an enhanced detail fatigue rating (DFR) method that is applicable to fatigue life prediction of TC4 bolted structures under thermal environments not exceeding 400 °C. Firstly, fatigue life data were acquired from base material specimens of TC4 titanium alloy tested at 20 °C, 200 °C, and 400 °C. Secondly, an enhanced DFR method that considered the temperature-dependent thermal influence was established based on the experimental results. The enhanced DFR approach was then applied to predict the fatigue life of double-shear TC4 bolted structures, and the results were compared with those obtained via the conventional DFR method. The findings demonstrate that the enhanced DFR method improves the average fatigue life estimation accuracy by 9.29% over the conventional DFR method within the 20~400 °C range. This establishes the proposed model as a highly promising tool for evaluating the fatigue performance of TC4 bolted structures under elevated thermal conditions below 400 °C.

Aircraft composite structures often feature stringers terminating near spars or cutouts, creating abrupt stiffness changes and stress concentrations prone to skin-stringer debonding under compression. This study investigates the failure characteristics and damage evolution mechanisms of stiffened composite panels with stiffener runout (SCP-SRs) from a thermodynamic perspective, based on stressing state theory. First, compression experiments were conducted on SCP-SR specimens to obtain their strain responses and identify failure modes. Subsequently, the strain data were modeled as state variables of the structure. Using the network-free renormalization approach, we extracted modal features and characteristic parameters that describe the stressing evolution process of the structure. By integrating the clustering algorithm, the phase transition loads at both partial and overall scales of the structure were identified. The stability of these phase transition loads was further verified using Wilson's theory. The results show that the standard deviations of the phase transition loads at the points of elastoplastic branching (EPB), failure starting (FS), and progressive failure (PF) were relatively low, with maximum values of 0.306, 0.458, and 0.416, respectively, indicating good repeatability and robustness of the method. The theoretical framework can provide a new engineering reference for strength design and failure analysis of large composite structures.

The present study addresses the challenges of stiffness and strength design for large-opening structures in special configuration aircraft such as blended wing-body aircraft. By combining mechanical principles and numerical simulation methods, this research investigates the stability analysis of long-span opening load-bearing beams. A method for analyzing the buckling characteristics of opening beams under the influence of multiple elastic supports from frames is established, providing design requirements for frame stiffness. Additionally, the study explores the collaborative optimization of structural arrangement and stiffness for opening beams and frames. Through case studies, the interaction between frame spacing and cross-sectional parameters of beams and frames is revealed. Furthermore, the research conducts failure analysis and load-bearing capacity prediction for aircraft opening structures, developing a coupled failure analysis method using a Nastran-Abaqus based multi-level non-intrusive model coupling analysis. Experimental validation is performed using typical large-opening structure failure tests. Comparative results show that finite element simulation effectively captures the nonlinear load and deformation behavior during testing, with maximum errors of 8.8% and 7.1% in load and deformation predictions, respectively. The simulation results also demonstrate good agreement with experimental buckling phenomena and failure modes.

Aircraft structural health monitoring technology plays an important role in the process of aircraft structur-al design,flight and maintenance.This technology can be used to predict structural health conditions and assist in maintenance and repair decisions.In this paper,several concepts of structural health monitoring and their applica-tions are introduced.The requirements of structural health monitoring are discussed,and the typical engineering cases of aircraft structural health monitoring technology are analyzed with the examples of F-35 and A400M.Then,the individual aircraft tracking and life control of typical aircraft,the load spectrum survey and life prediction of an aging aircraft are given,and the research on crack monitoring method and the main limitation of its application are discussed.The main idea of aircraft structural health monitoring system design and basic flow path of life predic-tion are put forward,and main tasks of structural health monitoring including selection of control points and flight parameters,construction of load/strain equation,calculation of damage and fatigue life,results output and verifica-tion of the equation are also introduced.Finally,the future research of aircraft structural health monitoring is pros-pected.

To improve the level of intelligent and information-based design of aircraft structural platform, the key link is to establish a reliable and effective monitoring system. Reasonable selection and optimal arrangement of sensor monitoring points are the key problems to be solved in the establishment of structural health monitoring system. This study is based on a wing box, the research route of monitoring point optimization based on load inversion is given, the correlation coefficient method is used to filter the monitoring points and eliminate a large number of closely related monitoring points, reduction of 996 monitoring points to 13 monitoring points, and the error difference between the actual loading load and the inversion load is obtained by multi-linear regression, the maximum error within 3

Operational loads of an aircraft are the prerequisite for assessing its safety or fatigue life. Traditionally, numerous strain gauge sensors are installed to monitor the operational loads, which inevitably increase the weight and system complexity of the aircraft. Therefore, in order to decrease the maintenance costs and data redundancy, the number and location of strain sensors should be optimized for accurate and reliable operational load monitoring. In this paper, a novel two-stage strain gauge location optimization method is proposed to reduce the number of strain gauges while maintaining the operational load monitoring accuracy, which is validated by a numerical case study of an aircraft wing. In the first stage, the traditional Pearson correlation measure is harnessed to initially eliminate numerous correlated strain gauge monitoring points, reducing 996 original strain gauge measurement points to 13 for the aircraft wing box. In the second stage, an improved correlation measure method is proposed to further reduce the 13 strain gauge points to 2, which can evaluate the correlation degree of several variables and simultaneously determine the optimal strain monitoring locations for the two load actuators in this study. The relative errors between the predicted loads and the actual loads for both load actuators are less than 4% when only two optimized monitoring points are adopted. In addition, a comparison study with LASSO regression and principal component regression methods is conducted. The results demonstrate that the proposed method has the characteristics of less monitoring points and higher load prediction precision.

The high overload and high maneuverability in the training course result in the structural damage and the life consumption speed is accelerated, the use of each model is obviously higher than the expected use, which greatly increases the risk of the operation of aircraft. The life design and verification work based on the design load spectrum in the design stage cannot meet the requirements of ensuring the flight safety in the whole life of aircraft. How to ensure the security of active military aircraft under the change of actual use mission has become an urgent problem to be solved. Based on the analysis method of structural fatigue individual aircraft tracking, this paper expounds the design method and task profile of individual aircraft, studies the calculation of equivalent damage of individual aircraft, puts forward three classification methods of usage severity of aircraft, which are mild severity, severity moderate and heavy severity, and carries out the determination of reference equivalent damage for the object aircraft. In addition, the interval classification method is used to classify the usage severity, and the method of the usage severity analysis is provided.

High-precision operational flight loads are essential for monitoring fatigue of individual aircraft and are usually determined by flight parameters. To tackle the nonlinear relationship between flight loads and flight parameters for more accurate prediction of flight loads, artificial neural networks have been widely studied. However, there are still two major problems, namely the training strategy and sensitivity analysis of the flight parameters. For the first problem, the gradient descent method is usually used, which is time-consuming and can easily converge to a local solution. To solve this problem, an extreme learning machine is proposed to determine the weights based on a Moore–Penrose generalized inverse. Moreover, a genetic algorithm method is proposed to optimize the weights between the input and hidden layers. For the second problem, a mean impact value (MIV) method is proposed to measure the sensitivity of the flight parameters, and the neuron number in the hidden layer is also optimized. Finally, based on the measured dataset of an aircraft, the proposed flight load prediction method is verified to be effective and efficient. In addition, a comparison is made with some well-known neural networks to demonstrate the advantages of the proposed method.

In this paper, some progress and achievements in aircraft integrity requirements, structural health monitoring, load spectrum measurement and life assessment research were presented. Several concepts of structural health monitoring were analyzed and compared, and the basic flow chart for health monitoring and life prediction of an aircraft structure was given. The selection of control points, construction of load/strain equations and stress calculation of control points were also described. Reliable IAT (Individual Aircraft Tracking) and life monitoring methods and software for IAT were developed for a certain type of aircraft, and fatigue life prediction of an aging aircraft was conducted based on actual measurement of load spectrum. The main features such as damage calculation, life evaluation and result output were discussed, and the future research focuses relating to intelligent structural health monitoring were finally explored.

多槽矩形开口结构是飞机上的一种特殊结构形式,在扭矩作用下的变形和应力分布状态复杂.文中推导了多槽矩型开口结构在扭转载荷下应力计算表达式,研究了其应力分布规律和特点.不需要借助有限元软件即可实现多槽型开口结构应力计算,在飞机设计初期可以加快飞机设计进程,提高计算效率;该文研究结果可用于飞机多舱体开口结构设计.

The aircraft structures combat survivability is one of the critical factors,which affect the integration combat efficiency of aircraft.The efficient design of aircraft structures survivability is of great significance to improve the combat ability,decrease the expenses in service,and ensure the well combat readiness of aircraft.At first,the design factors of aircraft structures combat survivability,based on the connotation of survivability and the design requirement of structures survivability for aircraft are combed.Then the design guidelines of aircraft structures combat survivability are discussed.Finally,the design and analysis methods of aircraft structures survivability are presented,under the effects of explosions-overpressure field and gust field,which are the two different aspects of blast wave.The analysis example is presented.Results show that the design factors of aircraft structures combat survivability are similar to the damage tolerance design,and the design process of damage tolerance could be used in structures combat survivability design.The fuselage panel,leading edge,moveable surface and cover,which are sensitive to overpressure of blast wave,can be designed to improve structure survivability by configuration optimization.

采用正交仿真试验法分析了整体壁板长桁端部斜削结构的斜削角、斜削角底部圆角半径、腹板根部圆角半径三个关键几何参数对细节应力集中系数的影响程度,并确定了影响该处细节疲劳性能的主要几何因素.开展了三长桁整体壁板疲劳试验,得到了长桁端部斜削结构的破坏寿命.采用细节疲劳额定值(DFR)法和名义应力法分别计算了该种结构细节的疲劳寿命.计算结果与三长桁整体壁板疲劳试验结果相比,疲劳危险部位一致,计算结果与试验结果吻合较好,且DFR方法预测结果更接近试验结果.

为了研究飞机复合材料T型加筋壁板结构在轴压载荷下的承载能力,对复合材料T型加筋壁板进行了轴压试验,并使用工程算法和有限元法仿真进行了分析,得到了加筋板的屈曲载荷及破坏载荷、载荷-位移曲线及损伤演化过程.对比两种分析方法与试验结果可以得到:当复合材料加筋壁板屈曲后,还有较强的后屈曲承载能力;相对于有限元分析方法,工程算法误差更大;准静态法可以有效地模拟出加筋板的屈曲及后屈曲行为,但是计算代价相对较大.

In this paper, residual compressive strength and compression-compression fatigue tests of two kinds of AS4/PEEK laminates with low-velocity impact damage were carried out to study the damage propagation characteristics. It was found that local buckling occurred around the impact dent and extended perpendicularly to the loading direction under fatigue compression loading. The expansion process of local buckling could be divided into three stages, and the damage mechanism of each stage was discussed. The experimental results showed that the impact damage and residual fatigue life both were quite dispersed under the same impact energy. A laminate with more severe impact damage had shorter residual compressive fatigue life, and the depth of impact dent could be used as an effective index to characterize the severity of impact damage and the residual fatigue life.

作为下一代大型民用飞机最具发展潜力的机型之一,翼身融合布局民机一直以来备受关注;而其非圆截面机身结构承载效率低和稳定性差等问题曾成为制约该机型发展的主要结构设计难题,由波音公司和NASA共同提出的基于PRSEUS的中央机体结构为解决这一难题提供了技术途径.为了深入研究翼身融合布局民机中央机体结构的承载特性、非线性分析方法及其优化策略,本文针对基于PRSEUS的HWB中央机体结构,依据其几何设计要求和复杂载荷特征,建立了高保真有限元模型及其边界、载荷施加方式;依据典型载荷工况,开展了HWB中央机体结构非线性分析;最后,利用多岛遗传优化算法,对HWB中央机体结构进行了方案优化设计.对比已有文献,结果表明:HWB中央机体结构非线性分析方法具有较好的变形分布预测能力;所提出的方案优化设计策略具有一定的工程应用价值.

拉挤杆缝合高效一体化结构(Pultruded rod stiffened efficient unitized structure,PRSEUS)综合利用了复合材料的一体化缝合和整体固化技术,能够满足翼身融合布局民机的传载、止裂、稳定性和维修性等结构设计要求.PRSEUS面板采用了低成本整体式结构设计制造方法,通过采用三维编织、单边缝合和可控气压树脂灌注等技术,完成纤维编织、缝合和树脂灌注过程,以低温共固化和缝合技术确保结构法向强度.通过合理选材和工艺设计,易于满足翼身融合布局民机不同部位结构的多样性设计需求.本文从翼身融合布局民机PRSEUS结构选材设计、PRSEUS结构制备核心工艺、PRSEUS工装夹具、典型PRSEUS测试壁板制造和典型机身试验件制造等方面系统阐述了翼身融合布局民机PRSEUS结构制造工艺技术的最新进展和发展现状,通过总结与展望PRSEUS结构制造工艺研究进展,为中国未来民机结构设计制造以及新型材料结构研发提供有价值的技术参考和研究方向.

本文介绍了飞机结构强度关键表征参数——应变监控技术的研究进展,探讨了当前基于飞行参数结合传统应变的结构部位受力状态监测技术的有效性,以及光纤光栅技术在结构应变监测中的应用.研究表明,采用飞行参数结合应变监控技术,可为飞机载荷监测与修正、结构响应监控、结构延寿和维修管理,提供重要的数据输入和技术手段.

The aircraft Structure Health Monitoring System (SHMS) is an important system for the old age or future aircrafts. The main components of this system includes many new and advanced sensors like fiber optic sensor etc. in the principal structure elements of aircraft structures, and the data processing module which needs to collect and demodulate the massive data, and the data analysis and storage which can be used to predict the structure’s life and evaluate the health status of the whole aircraft. This paper introduces the concept and functions of the SHMS, and presents the state-of-the-art for SHMS all-around of China, and the development history of SHM and aircraft life management during the past several decades. As the technology of big data analysis and cloud computing becomes more and more mature, a new research direction has been proposed in this paper. This paper provides the operating principle of SHMS and the flow of big data including the data collection and analysis and management. The system frame design method is proposed, and both of software and hardware conception also be described in this paper. The main functions and maintenance of SHMS also be discussed for the system design. Last but not the least, this paper provides the assumption that the aircraft structure life follows the lognormal distribution and the research on comparison between the aircraft structure life with and without structure health monitoring system. The analysis result shows that the aircraft use life and average life grows as the reliability of structure health monitoring grows and the aircraft structure maintenance become more effective.

本文研究了各种等损伤计算方法,推导了指数形式的等损伤计算公式对应的过载转换公式,并验证了与其他等损伤计算公式的关系.以典型飞机疲劳试验载荷谱为基础,结合各种等损伤计算公式,通过数值计算确定了基于重心过载和单机寿命监控的当量损伤计算公式及损伤指数.在此基础上,分析确定了单机寿命监控的采样率和滤波门槛值,为实施单机寿命监控提供技术支持.

由NASA和波音公司共同提出的拉挤杆缝合高效一体化结构(PRSEUS),由于具有优异的抗压稳定性和止损/止裂等承载优势,已成为解决翼身融合布局民机非圆截面机身结构承载效率低和稳定性差等问题的主要途径.本文针对典型PRSEUS受压壁板结构,开展了线性/非线性屈曲及渐进损伤分析;提出了综合考虑蒙皮、止裂带、长桁翻边、隔框翻边等一体化缝合元件贯穿支撑构型几何关系和偏置参考面的建模方法,提高了PRSEUS受压壁板有限元模型的精度;提出了综合考虑屈曲特征值、非线性屈曲载荷等多影响因素的网格收敛性分析方法,提高了PRSEUS受压壁板屈曲分析的计算效率;提出了最小屈曲特征值、几何节点偏移以及最小屈曲特征值-几何节点偏移组合式等3种初始缺陷引入方法,提高了PRSEUS受压壁板损伤分析的计算精度;完成了基于纤维与基体损伤本构关系的典型PRSEUS受压壁板非线性屈曲损伤分析,通过与试验结果对比,给出了针对PRSEUS结构的非线性屈曲渐进损伤演化分析方法.为翼身融合布局民机PRSEUS结构的稳定性/损伤分析和设计提供了方法和技术支撑.