基于应变能法的单搭接螺栓剪切模型

为了简化复杂结构在冲击数值分析中的大量螺栓连接,可用等效的载荷位移模型代替复杂的螺栓连接关系,本文中针对单搭接螺栓连接在剪切载荷下建立了连接本构关系。首先通过对有预紧力的单搭接螺栓进行实验和精细有限元模拟,揭示了螺栓剪切载荷位移曲线的特征并针对不同特征阶段进行了相应的物理机理分析。在此基础上对于载荷位移曲线的界面黏结、部分滑移、整体滑移阶段提出了连接本构模型的基本形式和各阶段的参数估算方法。在部分滑移阶段考虑了4个方面的刚度贡献,其中部件对螺栓的支撑刚度是三维非轴对称变形问题,理论求解非常困难,本文中通过应力分布研究,采用应变能法解决了螺栓的支撑刚度的估算问题。提出的单搭接螺栓剪切模型物理含义明确,参数估算简单,准确度高。     

球形爆炸容器法兰联接螺栓的应变增长现象

本文在球形爆炸容器法兰连接螺栓的动态响应实验研究中,发现了法兰螺栓的应变增长现象。以实验数据为基础,通过数值模拟对螺栓发生应变增长的原因进行分析。研究结果表明:受到开口结构的影响,端盖螺栓受到的压力载荷强度较容器内壁大幅增强。螺栓的应变增长是被增强压力载荷的前2个周期与结构响应形成共振时引发的,通过改变端盖质量、预紧力大小可以避免螺栓发生应变增长。     

扭转剪应力对螺栓紧固应力声弹性测量的影响研究

1 引言螺栓作为一种紧固件,广泛应用于桥梁、重型机械、大型密封装置、输电塔等各种结构中.结构的整体性,可靠性,在很大程度上取决于螺栓紧固应力的大小是否合适.因此,确定和监测螺栓的紧固应力就显得至关重要.传统的测量方法是采用专用扭力板手,经扭矩推算出螺栓轴力.但由于螺纹面之间及螺母与工件之间摩擦系数这个不定因素,其测量精度受到限制.且这种方法不适用于后期监测.超声法—利用声弹性原理直接测量轴力,精度较高,且便于现场应用和长期监测.但这种新方法在现场应用尚有些问题需要研究.其中包括扭转剪应力对声弹性应力测量的影响.已紧固的螺栓,除了受轴力外,还受     

缝合复合材料泡沫夹层结构层间剪切性能研究

复合材料泡沫夹层结构在飞机结构上的应用日益广泛,面板与芯子的全厚度缝合是提高泡沫夹层结构层间剪切性能的重要途径。本文通过层间剪切试验,研究不同缝合密度下单向带面板和织物面板复合材料泡沫夹层结构的层间剪切强度及其变化规律。研究结果表明:面板为单向带的缝合泡沫夹层结构最大剪切载荷和层间剪切强度均优于未缝合泡沫夹层结构的,且层间剪切性能随着缝合密度的增加而增加;面板为织物的缝合泡沫夹层结构缝合密度超过4000针/m2左右后,最大剪切载荷和层间剪切强度均优于未缝合泡沫夹层结构,且层间剪切性能随着缝合密度的增加而增加。     

水陆两栖飞机尾翼抗鸟撞设计与验证方法研究

为确保水陆两栖飞机尾翼结构的抗鸟撞性能,针对其不同结构部位提出不同的抗鸟撞设计思路。耦合SPH方法建立了尾翼结构的鸟撞数值模型,采用实验方法获得了结构铝合金材料的准静态和中低应变率拉伸实验数据以及不同冲击速度下带母材铆钉的极限拉伸载荷和极限剪切载荷数据。进一步开展了尾翼结构抗鸟撞分析,并通过鸟撞实验对数值分析结果进行验证。结果表明,针对水陆两栖飞机尾翼前缘结构提出的两种抗鸟撞设计思路合理,且具有较好的抗鸟撞性能;结构采用的3种铝合金存在较为明显的应变硬化效应,但应变率敏感性较弱;随着加载速度的增大,结构采用的4种铆钉拉伸载荷呈下降趋势,但总体幅度并不大,而剪切载荷变动较小;建立的尾翼结构鸟撞数值分析模型准确,较好预测了结构的破坏模式和鸟体冲击分散过程。     

平纹编织陶瓷基复合材料面内剪切细观损伤行为研究

采用约西佩斯库(Iosipescu)纯剪切试件,研究了平纹编织SiC/SiC和C/SiC复合材料的面内剪切应力-应变行为和细观损伤特性.通过试验获得了材料不同方向上的单调和迟滞应力-应变行为,对比分析了两种材料的剪切损伤特性,结果表明材料的剪切损伤演化规律受热残余应力水平影响严重.由试件断口电镜扫描结果发现剪切加载状态下桥连纤维承受显著的弯曲载荷和变形,据此提出了纤维弯曲承载机制,并结合裂纹闭合效应分阶段阐释了材料的剪切迟滞环形状.基于材料的剪切细观损伤机制,通过两个损伤变量表征了材料的剪切损伤演化进程,得到了材料的面内剪切细观损伤演化模型.对比发现2D-C/SiC复合材料45°方向基体裂纹的起裂应力明显小于2D-SiC/SiC复合材料,而两者0°/90°方向裂纹的起裂应力基本相同.      

CFRP加筋板剪切稳定性试验和数值分析模型研究

为研究数值模型对碳纤维增强复合材料(CFRP)加筋板面内剪切稳定性试验计算结果的影响,采用画框式夹具对CFRP加筋板进行了屈曲试验,获得了初始屈曲载荷和载荷-应变曲线;同时基于ABAQUS有限元软件建立了四种数值模型,分别进行线性屈曲和非线性屈曲分析,通过将计算结果与试验结果的比较,确定了有效的数值模型．在此基础上,通过对屈曲前、后有效区边界上的内力分布的比较分析,明确了夹具的传力效果、试件真实的受力状态,以及偏差产生原因．     

垫衬补偿对含翘曲间隙L型层合板力学性能的影响研究

复合材料结构整体化制造中极易产生由固化变形引起的局部翘曲，该翘曲会显著削弱螺栓装配下复合材料结构的承载能力。通过试验和数值分析开展了垫衬补偿对含翘曲间隙L型层合板极限载荷恢复效率以及失效行为的影响研究。通过微观CT表征了含翘曲间隙L型层合板在螺栓装配后的损伤分布特征，并测试了其极限承载能力；对比分析了垫衬补偿对螺栓装配后含翘曲间隙L型层合板损伤和极限承载能力的影响；并借助数值分析手段定量地研究垫衬补偿对含翘曲间隙L型层合板承载能力恢复的影响机制。结果表明：紧固强制消除翘曲间隙会在螺栓装配区域和拐角处出现一定程度的分层损伤，其中拐角区域的分层损伤可能是促使加载过程分层快速扩展，并导致极限承载能力大幅下降的主要原因；垫衬补偿技术极大的减小了装配损伤的出现，尤其是避免了拐角区域分层损伤的快速扩展，有效提高含翘曲间隙复合材料结构的极限承载能力。     

六自由度静定支持与约束技术在飞机载荷校准试验中的应用

在研究以往飞机载荷校准试验时非静定支持与约束方式的基础上,首次提出了飞机六自由度静定支持与约束方法;根据载荷校准试验得到的飞机受力分析结果,对飞机起落架支持与约束载荷进行了分解和简化,建立了飞机六自由度静定支持与约束工程模型;结合某型飞机的载荷校准试验对该方法进行了试验验证,结果表明:采用该方法可实现对飞机约束载荷的理论计算和实时监控,保证了飞机机翼等复杂结构部件在高载校准工况下试验过程的安全、可控;使飞机机翼的试验载荷量级提高到了限制载荷的43%,机翼载荷建模精度可控制在3%以内。     

弹塑性复合材料力学性能的细观研究

应用细观力学的Eshelby等效夹杂理论研究了复合材料的弹塑性问题。以铝基复合材料为例,建立了多轴载荷下复合材料弹塑性应力-应变关系,并且理论预报与实验结果符合较好,分析了夹杂形状、体积分数及加载路径对材料宏观性能的影响。同时,还研究了热塑性复合材料热膨胀系数与工艺温度之间的变化规律,分析了热残余应变对材料设计的影响。     

2012 William M. Murray Lecture: Some Curious Unresolved Problems, Speculations, and Advances in Mechanical Fastening

Mechanical joining is one of the oldest, most important, and most neglected aspects of engineering design of machines and structures of all types and sizes. Approximately 250 U.S. companies manufacture fasteners worth over $8 billion per year. There are 18,000 fasteners in a common fighter jet airframe, and fasteners account for roughly one-third the cost of a typical airplane. Yet, most failures of structures, including aircraft, originate at fasteners, suggesting that improved understanding of fastener mechanics, better design criteria, and informed applications of fundamental knowledge are required. This issue is exacerbated by increased demands on systems, particularly in the transportation and military sectors, and by the growing use of composites, for which current fastening practice seems to be underdeveloped owing to the complexities of material structure and response. This lecture first traces a brief history of mechanical joining, its importance, and the problems faced by engineers in designing for fastening. Research and development of fasteners through analysis and experiment are complicated by the large array of variables involved, and investigators must have at hand an extensive array of experimental and analytical techniques as well as an appreciation of the practicalities of fastening. Verification and validation of findings are crucial, and extrapolation is fraught with pitfalls. Described subsequently are some examples of experimental results that generate speculation and that might provide points of entry for investigators who are willing to take on difficult challenges where progress would be valuable but is not easily realized. These cases include, among others, odd and perhaps dangerous behaviors resulting from coldworking holes in engines and structures, impact stresses caused by joint slippage in composites, the use of inserts to control stress concentrations, difficulties in applying sufficient clamping force in composites, merits of hole shaping, and unusual configurations of conical washers. Finally, some ideas for hybrid joining and for systems that allow quick field assembly/disassembly are briefly described.     

复合材料胶接修补问题的试验研究和分析

用带中心裂纹的Ｌｙ１２ＣＺ铝合金板模拟飞机的损伤结构,对于不同的修补方式（单边和双边修补）,不同的补片材料（铝合金和复合材料）和几何尺寸以及脱胶等因素,进行了静强度和疲劳裂纹扩展试验． 实验表明． 胶结修补能明显地提高损伤结构的强度和疲劳寿命,同时实验还为修补实践的选材和几何参数的设计等提供了大量数据． 文中最后给出了修补结构疲劳裂纹的有限元计算估计,并与试验结果进行了比较和分析．     

An instrumented fastener for shear force measurements in joints

A preliminary investigation has been conducted on instrumented fasteners for use as sensors to measure the shear loads transmitted by individual fasteners installed in double-splice joints. Calibration and load verification tests were conducted for instrumented fasteners installed at three fastener torque levels. Results from calibration tests show that the shear strains obtained from the instrumented fasteners vary linearly with the applied load and that the instrumented fasteners can be effectively used to measure shear loads transmitted by individual fasteners installed in double-splice joints. Tests were also conducted with three instrumented fasteners installed in a typical double-splice joint. The test results showed that the load distribution between individual fasteners is dependent on the location of the fastener in the joint and the fastener torque level. The fastener located near the end of the joint with the single plate carried more load than the fasteners located near the end of the joint with the two plates. Installing the fasteners with a torque greater than finger tight results in a significant amount of the load being carried by friction between the faying surfaces of the plates even if the faying surfaces are polished and lubricated. Increasing the fastener torque increases the load being carried by friction between the faying surfaces of the joint. Increasing friction between the faying surfaces of the joint. Increasing the fastener torque also results in a more uniform distribution of the loads between the individual fasteners for joints in aluminum plates with two fasteners, but does not have a significant effect for joints in steel plates with three fasteners. Paper was presented at the 1992 SEM Spring Conference on Experimental Mechanics held in Las Vegas, NV on June 8–11.     

电磁加载下7075铝环的膨胀断裂模式转变研究

利用电磁膨胀环实验技术,研究了7075铝环在2700～8700 s?1拉伸加载应变率下的断裂模式转变现象。实验结果表明:在低应变率下,铝环的断裂受最大正应力控制,发生拉伸断裂;在高应变率下,铝环的断裂受最大剪应力控制,发生剪切断裂;在中应变率下,铝环的断裂同时受最大正应力和最大剪应力控制,为拉伸和剪切同时存在的拉剪混合断裂模式;随着应变率的增加,铝环的破片数量呈先增加、再减小、最后增加的趋势,破片数量变化拐点与断裂模式转变点基本吻合。     

极端环境下连续碳纤维增韧的陶瓷基复合材料的力学行为

连续纤维增韧的碳化硅复合材料(以下简称C/SiC),作为超高速飞行器热结构使用时,有可能在高温环境下受到高速撞击的作用,因此,掌握其在极端环境(高温、高应变率)下的力学性能是进行结构安全设计的基础。本文采用具有高温实验能力的分离式Hopkinson杆,在293~1273K温度范围内进行了动态压缩力学性能测试,研究了环境温度和加载速率对材料力学性能的影响。结果表明:C/SiC复合材料的高温压缩力学性能主要受应力氧化损伤和残余应力的共同影响。实验温度低于873K时,应力氧化损伤的影响很小,而由于增强纤维和基体界面残余应力的释放使界面结合强度增大,复合材料的压缩强度随温度的升高而增大;当实验温度高于873K时,应力氧化损伤加剧,其对压缩强度的削弱超过残余应力释放对强度的贡献,材料的压缩强度随温度的升高逐渐降低。由于应力氧化损伤受应变率的影响很大,当温度由873K升高至1273K时,高应变率下压缩强度降低的程度要比应变率为0.0001/s时低得多。     

<Emphasis Type="Italic">In-Situ</Emphasis> CT Damage Analysis of Metal Inserts Embedded in Carbon Fiber-Reinforced Plastics

Carbon-fiber-reinforced plastics (CFRPs) are gaining increasing applicability to lightweight structures (e.g., automotive applications) due to their outstanding mechanical properties. High-performance parts can be fabricated from CFRPs, but they have the disadvantages of low shear and bearing strength. To achieve detachable connections and introduce loads without decreasing the load-bearing capacity of the composite, it is important to use mechanical fasteners without drilling into the parts. To accomplish this, metal elements called inserts are embedded in the CFRP laminate. Damage behavior in a CFRP under tensile conditions has several different mechanisms, depending primarily on the deformation of the insert. This research investigates the in-situ failure behavior of the composite under tensile loads by investigating the deformation of the insert via computed tomography (CT). The results are also used for validation of the insert’s deformation using a finite-element model (FEM).     

Strength and stiffness of sandwich beams in bending

This investigation is concerned with the experimental versus analytical correlation of the mechanical properties of sandwich-beam specimens. Such sandwich structures are commonly employed in the aircraft industry. Four-point and three-point load tests were conducted on a large number of sandwich-beam specimens, fabricated by using fiber-glass reinforced plastics (both unidirectional and woven-glass cloth) and DTD 685 aluminum alloy for the facings with aluminum honeycomb core and polyurethane foam cores and the indigenously available Araldite as the bonding medium between the core and the facings.The flexural stiffness of the composite sandwich specimens used in this investigation compared favorably with theoretical predictions. The shear stiffness was found to be about 55 percent and 45 percent of the theoretically predicted values for FRP (fiberglass-reinforced-plastic) cloth and FRP unidirectional laminates with aluminum honeycomb core sandwich, respectively. The failure load as determined by experiments was less than the theoretically predicted safe load. There was a loss of strength as well as a steep decrease in the failure load in the case of low density foam core.It was concluded that FRP facing plates with aluminum honeycomb core sandwich structure may be preferred to similar aluminum-alloy facing sandwich construction if high flexural stiffness and shear stiffness properties are required at less cost and weight. Indigenously available Araldite was quite satisfactory for bonding the core to the facings.This investigation has confirmed the importance of experiments in the field of sandwich structures which can effectively replace other conventional uneconomical structural or machine members which are currently in use.     

石英纤维布氰酸酯树脂基复合材料的环境试验性能研究

对石英纤维布增强氰酸酯树脂基复合材料进行了常温、低温和湿热三种环境下的拉伸、压缩、面内剪切、层间剪切和钉孔挤压等试验研究,得到了该复合材料的拉伸压缩面内剪切的强度和弹性模量以及层间剪切强度和钉孔挤压强度等力学性能参数。结果表明:石英纤维布增强氰酸酯树脂基复合材料的力学性能参数具有温度相关性。在低温环境下,力学性能参数不同程度地增强,最大增长量达40%;而在高温环境下,力学性能参数则明显下降,最高下降量达到56%。另外,由于复合材料的层间性能主要由基体决定,所以湿热环境对复合材料层间性能的影响很大,在工程实际中应特别关注。     

复合材料层板构件气弹振动控制实验研究

飞机局部复合材料构件存在着因涡流作用诱发的弹性振动问题。本文首先研究了表面粘贴型压电元件对复合材料构件的传感和驱动原理;针对结构待控模态的要求,采用了D-准则优化确定同位压电传感/驱动器在构件中的布置方案;应用自适应振动前馈控制原理和方法,构造了闭环控制系统;分别利用正弦和方波信号激发气动声场,对玻纤/环氧圆筒构件进行激振和振动控制实验;结果表明：构件主要待控模态的振动得到了有效抑制,但也出现了高阶模态被激发的问题,导致结构辐射噪声上升。     

Composite patch reinforcement of cracked aircraft upper longeron: analysis and specimen simulation

The use of composite patches on cracked portions of metallic aircraft structures is an accepted means of improving fatigue life and attaining high structural efficiency. As more and more advanced composite materials are beng developed, the wider use of the repair technology is anticipated even for the reinforcement of primary aircraft structure. The objective of this work is to illustrate how the composite patch repair technology can be successfully applied to restore the structural integrity of cracked components.The Phosphoric Acid Anodize (PAA) surface treatment on aluminum when applied in conjunction with the AVI13/HV998 adhesive were essential for achieving the appropriate patch bonding strength. Such a process was done without immersing the component into the PAA tank; dismantling the component from the aircraft was not necessary. Boron/epoxy and carbon/epoxy patches were applied at room temperature to the 7075-T6511 cracked specimens and tested under fatigue simulating the load spectrum for the upper longeron attached to the access door of the electronic equipment bay. Considerable improvement in the fatigue life was observed after the repair. Equivalent flight test hours were increased from approximately two thousand hours at which the component fractured completely when not repired to twelve thousand hours when the repair was made with only a small amount of crack growth. A six times increase fatigue life is obtained. The laboratory developed technique has been applied to several in-service aircraft which have now been flown for more than 700 h without detection of crack growth.