Investigation of a Boeing 747 wing main landing gear trunnion failure

A loud noise was heard from the vicinity of the port wing landing gear during pushback of a Boeing 747-300 from the terminal at Sydney (Australia) airport. Inspection showed that one of the wing landing gear trunnion fork assemblies had failed. Detailed investigation revealed that the trunnion had failed by fatigue cracking. Deep machining grooves were found at the root of an internal radius that had not been shot-peened as required, and a chemical surface process during manufacture had resulted in shallow intergranular attack at the bottom of these grooves. It is probable that the critical cracking started from some of these grooves. In addition, the wall thickness at the failure location was significantly less than the minimum required in the drawings.Since the deep machining grooves, the lack of peening and the intergranular attack were all consequences of manufacturing, the fatigue cracking probably started shortly after the component entered service. This implies that fatigue cracking was present during all the trunnion overhauls, but was not detected by non-destructive inspections during the overhauls. Quantitative fractography was used to produce a crack growth curve based on fracture surface markings thought to represent the overhaul timings. The crack growth curve suggested that the fatigue cracking was large enough to be detected by inspection during the last overhaul, if not the one before. However, it was probably not easy to detect the cracking. This investigation therefore highlights the difficulties that can be encountered when inspection is the last (or only) line of defence against failure owing to unexpected manufacturing deficiencies.     

Failure analysis of nose landing gear assembly

This paper analyses the causes of the incident of an F-4 which landed abnormally due to a crack in the nose landing gear hydraulic actuator cylinder. Striations, typical of a fatigue crack, were observed on the fracture surface and a large quantity of corrosive oxide was detected in the initial stage of the crack. Furthermore, surface pits caused by corrosion were found inside the cylinder. A surface flaw of this kind was connected to the fracture surface as a microscopic crack. By analyzing the fractography and metallography of the failed parts, it is found that the cylinder pitted as a result of chemical reaction with internal fluids, the pitting became a microscopic crack and caused failure.     

基于虚拟样机技术的飞机起落架着陆载荷分析

起落架是飞机重要的承载机构,其在飞机的起飞、降落和着陆过程中担负着十分重要的作用。起落架着陆载荷是指在飞机着陆瞬间起落架所承受的地面载荷,根据位置的不同可分为轮胎接地点载荷、轮轴点载荷和交点载荷。飞机设计重量（包括空机重量、商载和燃油重量）的分布变化会对起落架着陆载荷产生影响。以典型民用飞机的起落架为研究对象,基于虚拟样机技术对起落架模型进行着陆分析。首先利用HyperMesh、MSC.Nastran软件对飞机起落架模型进行前处理和简化,然后根据中国民用航空规章运输类飞机适航标准相关条目要求,在MSC.ADAMS软件中设置相应的工况参数并进行起落架着陆动力学仿真分析,研究不同燃油密度下飞机的重量和重心变化对起落架轮轴点着陆载荷的影响。通过对比仿真结果可以发现：飞机的重量和重心变化对起落架轮轴点着陆载荷产生了明显的影响,而燃油密度变化对起落架轮轴点着陆载荷的影响较弱。研究结果可为今后有针对性地开展起落架着陆载荷计算提供参考和指导。     

Failure analysis of a nose landing gear fork

This paper presents a detailed analysis of a nose landing gear failure. The developed study comes following an accident occurred in which the nose of the landing gear's fork of a light aircraft failed during landing. According to Federal Aviation Administration, in average, 55% of aircraft failures occur during takeoff and landing.In order to determine the causes of the accident, a material analysis was performed, followed by a detailed study of the fracture's surface both visually and using optical and scanning electron microscopies. It was observed that fatigue cracks developed in the vicinity of the bolted holes, which work as supporting connections, on the topside of the nose fork and, as such, it can be concluded that the referred area was subjected to cyclic stresses originating and propagating cracks in the material. This cracking is characteristic of the existence of stress concentration areas. Identified the crack initiation zone with ratcheting and beach marks near the origin of the crack, combined with the fact that the nose wheel fork is subject to cyclic loading, leads to the conclusion that the component failed due to fatigue.Finite element analyses were also performed on the nose fork taking into account service conditions in order to assess the structural integrity of the component. During the analysis it was observed that the critical areas are located in the vicinity of the connecting holes, as it was observed in the fracture surface analysis. The assembly behavior in the presence of four straight cracks, originating from the fork holes, was also studied using the stress intensity factors, calculated using the contour integral method.     

Head-up displays assist helicopter pilots landing in degraded visual environments

Civilian rotary-wing aircraft pilots typically rely on visual information from the external environment to guide flight, but are increasingly required to operate in degraded visual environments. The current study evaluated the impact of a head-up display (HUD) upon pilot performance, perceived situation awareness and workload. A 2 × 2 repeated measures design required qualified rotary-wing pilots (N = 6) to fly in clear and degraded visual conditions both with and without the HUD. In degraded visual conditions, the HUD significantly improved pilot perceived situation awareness whilst simultaneously reducing workload. Objective flight technical performance data offered preliminary support for a positive change in pilot behaviour when using the HUD in degraded visual conditions.     

Investigation on the failure of the gear shaft connected to extruder

In respect to the fracture of the gear shaft of an extruder, the torque on the shaft replaced the fractured one is measured under typical work conditions. The fatigue and material mechanics properties of the fractured shaft are obtained by tensile and fatigue test. Based on the analysis of cyclic fatigue load and the stress of gear shaft, the reason of the shaft fracture is discussed. The analysis indicates that when the extruder is running, high mean stress occurs in the gear shaft, the failure of the gear shaft was caused by fatigue and material strength degradation. Finally, suggestions are given to guide the future operating of the extruder to prevent the gear shaft from fracture.     

起落架摆振主动控制分岔研究

研究起落架摆振及其主动控制的非线性动力学问题。根据分岔理论,应用数值延拓算法求解参数化非线性摆振微分方程,得到摆振的Hopf分岔点(临界摆振速度)和摆振极限环的幅值,通过李亚普诺夫指数确定极限环的稳定性。在此基础上,提出采用主动控制策略抑制摆振,建立相应数学模型,应用数值延拓算法求解加入控制后的非线性摆振微分方程。结果表明,适当选取控制参数,在飞机滑跑速度范围内,应用主动控制策略可以达到完全抑制摆振的目的,并且对系统参数的变化具有强鲁棒性。     

起落架载荷标定模型的集成学习方法

为解决飞机起落架载荷标定实验使用线性回归建立标定方程结果不理想的问题,考虑到实验中起落架压缩行程和应变片布片位置等因素对标定载荷的非线性影响,运用特征融合、集成学习理论,通过使用AdaBoost和XGBoost非线性回归方法,构建起落架载荷标定模型。首先,通过起落架载荷标定实验获取实验数据,使用主成分分析方法建立输入特征矩阵;其次,构建起落架载荷标定模型,将起落架三向加载载荷分别作为标签向量,训练集和测试集根据随机取样原则划分,使用AdaBoost和XGBoost两种方法训练标定模型;最后,在测试集中对载荷进行拟合预测,并使用均方根误差、平均绝对误差、决定系数、耗时4个评价指标对模型进行评估。实验结果显示,与广泛使用的最小二乘法相比,XGBoost方法建立的标定模型能够更好地拟合加载载荷,在不考虑时效性的场景下XGBoost算法更具优势。研究结果对提高飞机起落架载荷实测准确性以及飞机结构健康监测的进一步研究具有重要价值。     

基于可变形空腔的起落架舱体噪声抑制研究EI北大核心CSCD

提出了一种基于可变形空腔的起落架舱体结构,通过机械装置调节舱体底板及后壁倾斜角度,不需要额外增加舱体体积,使用声学有限元法探讨了该结构在低马赫数下的噪声抑制效果。研究发现:随着舱体后壁倾斜角度的增大,舱体内部及外部的噪声明显减小,同时模态频率逐渐增大,有助于避免舱体结构发生共振破坏;舱体后壁倾斜一个较小的角度就能有效地改善内部的声反射环境,进而抑制舱体内部的高频模态噪声、总声压级。当后壁倾斜角度大于某个临界值时,继续增大倾斜角度对于舱体内部高频模态噪声以及总声压级的抑制效果不再明显,在当前的仿真条件下,舱体后壁最佳倾斜角度范围为10°~16°。     

ARJ客机进场飞行过程的起落架噪声的实验研究

对ARJ客机降落阶段的起落架噪声进行了现场测量分析,为了将起落架噪声源与其他噪声源分开,利用了阵列信号处理技术,实验发现:ARJ飞机起落架噪声主要集中在900 Hz以下,且前、主起落架均存在强单音噪声,对起落架噪声可能的声源来源进行了分析,为起落架降噪设计提供了支持。     

The effect of overloading sequences on landing gear fatigue damage

In service, landing gear can be subject to unexpected hard landing load, which is beyond the design domain. The consequences due to overload can affect the design life of a landing gear to some extent. In this paper, the effect of overload and different loading sequences in random spectra on fatigue damage are investigated, using strain–life based fatigue analysis methods. The discussions are emphasised on the effect of loading sequences on residual stress and mean stress, especially the effect of overload on the fatigue damage of subsequent cycles. In addition, different fatigue analysis techniques in commercial fatigue analysis packages are reviewed and compared. The analysis indicates that the overload effect is stress state dependent and dominated by local residual and mean stress. A ‘Begin’ overload in a load spectrum would cause more damage in the local compressive yield area and an ‘End’ overload within a spectrum will worsen the tensile yielding area. It is suggested that the load sequence effect should be considered in common fatigue analysis if local yielding would exist before or after overloading.     

Failure analysis of an aircraft nose landing gear piston rod end

A failure investigation has been conducted on a piston rod end used in a hydraulic actuating cylinder of an aircraft landing gear. The failed piston rod end was found to be broken. An evaluation of the failed piston rod end was undertaken to assess its integrity that included a visual examination, photo documentation, chemical analysis, hardness measurement, tensile testing, and metallographic examination. The failure zones were examined with the help of a scanning electron microscope (SEM) equipped with EDX facility. A stress analysis is also carried out by the finite element technique for the determination of highly stressed regions on the piston rod end. The results indicated that the piston rod end failed by fatigue with cracks initiated at the surface close to the mechanically damaged region due to high stress concentrations.     

海鸥300飞机起落架着陆动力学试验研究

针对海鸥300水陆两栖飞机起落架落震试验的技术需求,研制一套用于轻型飞机起落架落震试验的动力冲击试验系统,解决了低摩擦滑道设计、机轮高速带转、跑道模拟、撞击平台设计等关键技术.结合CCAR-23-R3要求,完成了海鸥300前起落架落震调参试验、限制落震试验和储备能量吸收落震试验.结果表明,试验系统工作稳定可靠,起落架限...     

An equivalent uniaxial fatigue stress model for analyzing landing gear fuse pins

We study the biaxial stress state conditions in landing gear fuse pins in the fuse groove. This biaxial state comprises a combination of shear stresses which are usually the largest stresses in the fuse pin by design, and compressive stresses which keep the half-section of the fuse pin in equilibrium. Conventional fatigue analysis techniques use an equivalent uniaxial stress, based on the Mises stress of a pure-shear condition. The respective predicted fatigue damages are much higher than those obtained from fuse pin cyclic tests. A new equivalent uniaxial fatigue stress model is proposed that includes the additional compressive stress as a relief on the fatigue damage in the fuse groove, thereby explaining the observations from fuse pin tests. The model is used in conventional uniaxial strain-life fatigue software (Goodrich Aerospace’s Fatigue Life V2) to predict the fatigue damage on a landing gear fuse pin with a sample load spectrum. The results are then compared to the pure shear model, and to a biaxial finite element fatigue analysis. As compared to the equivalent Mises model, the proposed model provides less conservative estimation of the fuse pin fatigue life, the latter value being higher than that provided by the two-dimensional finite element calculation. __________ Translated from Problemy Prochnosti, No. 3, pp. 85–98, May–June, 2006.     

基于响应面方法的多支柱起落架着陆缓冲性能优化

基于起落架缓冲系统受力分析,在MSC.ADAM S平台上建立了某型飞机多支柱主起虚拟样机模型,在给定的投放质量和下沉速度下分析了该起落架的缓冲系统行程、轮胎垂直力和功量曲线。为降低飞机着陆载荷,选取对缓冲器性能敏感程度较大的参数为优化参数,包括空气腔初始容积、初始充气压力及主油孔面积,以着陆时轮胎最大垂直载荷为优化目标。通过基于正交设计的响应面方法,建立优化变量和优化目标间的数学关系,求解方法为模拟退火法。优化后,对新设计参数进行了虚拟试验验证,结果表明降低着陆总载荷8.3%,明显改善了起落架着陆缓冲性能。     

轮胎刚度智能预测及在起落架摆振分析中的应用

为分析飞机起落架摆振问题,需要获取航空轮胎的侧向、扭转刚度。建立了一套航空轮胎侧向与扭转刚度智能预测方法,并在某无人机起落架防摆设计中成功应用。收集了503组轮胎侧向、扭转刚度试验测试数据,来自3个轮胎厂商,7种轮胎型号,涵盖不同气压、载荷、压缩量、直径、宽度等特征参数的组合。建立了一个全连接人工神经网络模型,通过已有轮胎刚度测试数据对神经网络模型进行了训练,从随机分配的验证集训练效果可以看出,训练到1000次以后,均值误差基本收敛,预测值与真实值基本一致。利用训练后的模型对某无人机轮胎侧向与扭转刚度进行了预测,并将预测结果应用于某无人机起落架摆振稳定性分析,确定起落架临界阻尼,指导减摆器的设计。     

管簧式起落架落震试验及分析

为了获得管簧式起落架着陆性能的特点,设计了管簧式起落架落震试验,以获得相关的落震性能参数。同时设计了落震测试系统,并通过对起落架落震系统的测试,确定落震试验操作系统的可行性。通过改变管簧式起落架下落高度和投放质量以及机轮带转等参数,总结机轮所受的最大垂向力和起落架最大垂向位移的变化规律,得到管簧式起落架的相关落震性能参数,为起落架设计提供参考。     

冲击载荷对飞机起落架螺纹连接的影响

对近期民航飞机主起落架连杆插销螺栓失效原因进行分析,估计出飞机降落过程中插销螺栓承受的最大冲击载荷为698 MPa,采用Yamamoto方法分析计算得到插销螺栓每一个啮合螺纹的应力分布,利用有限元分析法建立了连杆插销螺栓连接的有限元模型,分析所有啮合螺纹的应力应变场。分析了冲击载荷对螺旋效果和螺纹啮合位置产生的影响,这为飞机起落架的校查及使用寿命预测提供了重要的依据。     

Tribological failure analysis of gear contacts of Exciter Sieve gear boxes

In the present paper the premature failure of gear contact encountered in Exciter Sieve gear boxes has been analyzed. The cause of gear contact failure is identified by simulating the load bearing capacity of lubricants and conducting controlled experiments on an Amsler disk-on-disk tribo tester. The results of performance behavior (i.e., load carrying capacity of lubricants, contact friction and weight loss of test specimens) of the simulated gear contacts have been reported. The theoretical and experimental results indicate presence of mixed to partial elastohydrodynamic lubrication conditions in the gear contact. To mitigate the problem of scuffing and scoring in the gear contacts, lubricating oils with extreme pressure additives and Base oil without additive have been tested and performance results are reported.