Journal articles on the topic 'Composite aircraft joints'
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1
Maier, Raluca. "Study on Increasing Performances of Hybrid Composite Through Pull Out Compression Test Assessement." Materiale Plastice 57, no. 1 (April 17, 2020): 329–35. http://dx.doi.org/10.37358/mp.20.1.5325.
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The exhaustive use of light advanced polymer composites into the development of the future high-efficiency long-range commercial aircrafts (primary) structures for increasing strength while decreasing weight for lower fuel consumption and less pollution has attracted great attention in the last decade. In this context, solutions for the integration and joining composites into the aircraft structures metallic sub-assemblies or components, while providing the necessary strength to sustain heavily loaded joints, are urgently required. The paper comes to meet an area of interest for many researchers and large integrators and manufacturers, by presenting several type of metallic join geometrical designs, along with metallic joint part integration solutions into a composite structure directly during production. Performances evaluation of the developed metallic join geometrical designs was performed in static regime by means of �pull-out� compressive tests. The paper focuses on the strength of the hybrid metallic/advanced polymer composites joints. It was concluded that an efficient design of the metallic joint and a proper transition from hybrid to the composite structure can lead to proper integration solutions while assuring safety requirements.
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Chowdhury, Nabil, Wing Kong Chiu, and John Wang. "Review on the Fatigue of Composite Hybrid Joints Used in Aircraft Structures." Advanced Materials Research 891-892 (March 2014): 1591–96. http://dx.doi.org/10.4028/www.scientific.net/amr.891-892.1591.
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The use of composite materials as a replacement for commonly used metals such as aluminium and steel are increasing in the engineering industry, particularly in the aerospace sector. The move towards light weight and high stiffness structures that have good fatigue durability and corrosion resistance has led to the rapid move from metal to composites. This change allows for further flexibility in design and fabrication of various components and joints. There are three main categories of joints used in composite materials – mechanically fastened joints, adhesively bonded joints and the combination of the two called hybrid joints. In order to adequately understand the effectiveness of these joints, substantial testing and validation is required, particularly in the use of hybrid joints for real life applications. Static testing, load distribution and parametric studies of hybrid joints have been investigated by various researchers; however further work is still required in understanding the durability and fatigue of hybrid joints and ensuring that both the adhesive and mechanical fasteners can work together effectively in producing an optimum joint. Mechanical fastening alone in composite laminates is not a preferred joining method as they create high stress concentrations around the fastener holes. Adhesive bonding although has numerous benefits it is difficult to detect the bond defect particularly in cases where weak bonds can occur during applications and it is sensitive towards the environmental conditions. Thus hybrid joints are seen arguably as being more effective in joining composite components together and offer greater residual strength. Hence the performance, strength and long-term durability of these joints need to be further investigated and be applied to practical situations whilst assisting in repair certification.
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Fortier, Vincent, Jean-E. Brunel, and Louis L Lebel. "Fastening composite structures using braided thermoplastic composite rivets." Journal of Composite Materials 54, no. 6 (August 14, 2019): 801–12. http://dx.doi.org/10.1177/0021998319867375.
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Aerospace composite material components are currently joined using heavy titanium bolts. This joining method is not ideal when considering its weight, thermal expansion, electrical conductivity, and risk of unbalanced load distribution. We propose here an innovative fastening technology using thermoplastic composite rivets. A rivet blank is heated above its melting temperature using Joule heating and is formed directly in the composite laminates by an automated process. Carbon fiber and polyamide blanks were used with two fiber architecture: 2D braid and unidirectional. The braided architecture showed superior manufacturing performance and repeatability. Joints were riveted in less than 40 s per rivet. The temperature measured in the riveted composite laminate in the vicinity of formed rivet reached only 136℃ during riveting. Double fastener lap shear testing showed breaking load of 6146 N per fastener. This joint strength is higher than comparable aluminum-riveted joints, and the specific joint strength is higher than titanium-bolted joints. With these advantages, the technology could be developed and used in the next generations of lighter, cleaner, and safer aircraft.
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Romano, Maria Grazia, Michele Guida, Francesco Marulo, Michela Giugliano Auricchio, and Salvatore Russo. "Characterization of Adhesives Bonding in Aircraft Structures." Materials 13, no. 21 (October 28, 2020): 4816. http://dx.doi.org/10.3390/ma13214816.
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Structural adhesives play an important role in aerospace manufacturing, since they provide fewer points of stress concentration compared to faster joints. The importance of adhesives in aerospace is increasing significantly because composites are being adopted to reduce weight and manufacturing costs. Furthermore, adhesive joints are also studied to determine the crashworthiness of airframe structure, where the main task for the adhesive is not to dissipate the impact energy, but to keep joint integrity so that the impact energy can be consumed by plastic work. Starting from an extensive campaign of experimental tests, a finite element model and a methodology are implemented to develop an accurate adhesive model in a single lap shear configuration. A single lap joint finite element model is built by MSC Apex, defining two specimens of composite material connected to each other by means of an adhesive; by the Digimat multi-scale modeling solution, the composite material is treated; and finally, by MSC’s Marc, the adhesive material is characterized as a cohesive applying the Cohesive Zone Modeling theory. The objective was to determine an appropriate methodology to predict interlaminar crack growth in composite laminates, defining the mixed mode traction separation law variability in function of the cohesive energy (Gc), the ratio between the shear strength τ and the tensile strength σ (β1), and the critical opening displacement υc.
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Rufin, Antonio C., and Robert L. Jochum. "Fatigue Analysis and Design of Composite-Metal Joints in Primary Aircraft Structures." Advanced Materials Research 891-892 (March 2014): 588–93. http://dx.doi.org/10.4028/www.scientific.net/amr.891-892.588.
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Increasingly broader application of composite materials in commercial airplane primary structure has led to a renewed, more in-depth assessment of the effect composite elements in mechanical joints have on the durability of the metal structure to which they are mated. Thermal stresses, clamp-up loss due to composite creep, composite joint properties, and lightning and electrical grounding effects have been found to be some of the parameters most significantly affecting or limiting the fatigue performance of the metallic constituents within these hybrid joints, influencing the way joint design and analysis are approached.
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McCarthy, Michael. "BOJCAS: bolted joints in composite aircraft structures." Air & Space Europe 3, no. 3-4 (May 2001): 139–42. http://dx.doi.org/10.1016/s1290-0958(01)90077-2.
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Chernov, Andrey, Danil Fomin, Ivan Kondakov, Ivan Mareskin, and Alexander Shanygin. "Lightweight and reliable metal–composite joints based on harmonization of strength properties of joined parts." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 232, no. 14 (May 24, 2018): 2663–72. http://dx.doi.org/10.1177/0954410018778797.
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One of the key problems of designing composite primary aircraft structures is the problem of development of lightweight and reliable joints for such structures. For conventional bolted joints used for connection of laminated composite primary structure elements, the loads are transferred via the contact of metallic and composite parts having considerably different strength and stiffness parameters. As the local strength properties of composites are defined by polymer resins, having several times less strength as compared to the metallic alloys, the maximal load that can be transferred via the joint is limited by the properties of resins. As a consequence, the metallic parts of such joints are loaded to a very low extent, that causes low weight efficiency. In the presented paper, the approach to development of metal–composite joints based on the principle of harmonization is proposed. The harmonization principle is to minimize the disbalance of strength properties of metallic and composite parts in the local zones of contact. This principle can be realized by two ways: either by removing the resin from the zone of contact to increase local stiffness of composite part (“stiff” joint concept) or by creation of metallic part with porous structure to decrease local stiffness of the metallic part (“soft” joint concept). In this paper, the evaluation of both concepts is given, based on numerical and experimental investigations carried out in this work.
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Balinova, Yu A., D. V. Graschenkov, A. A. Shavnev, V. G. Babashov, A. S. Chaynikova, E. I. Kurbatkina, and A. N. Bolshakov. "High-temperature heat-shielding, ceramic and ceramic-metal composite materials for new-generation aviation equipment." Journal of «Almaz – Antey» Air and Defence Corporation, no. 2 (July 19, 2020): 83–92. http://dx.doi.org/10.38013/2542-0542-2020-2-83-92.
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This paper describes achievements of the All-Russian Scientific Research Institute of Aviation Materials in the field of creating high-temperature heat-shielding, ceramic and metal-ceramic composite materials. The advantages and prospects of applying the developed materials in the manufacturing of structural elements of aircraft and friction joints are discussed. The synthesis features and basic properties of metal-ceramic composite materials based on light alloys, refractory metal matrices, ceramic composite materials for use in heavily loaded structural elements of modern aircraft are presented. The main achievements in the field of heat-shielding materials based on refractory oxide fibres are presented, along with their properties and application in new-generation aircrafts.
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Jadee, Khudhayer J., and A. R. Othman. "Fiber Reinforced Composite Structure with Bolted Joint – A Review." Key Engineering Materials 471-472 (February 2011): 939–44. http://dx.doi.org/10.4028/www.scientific.net/kem.471-472.939.
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Fiber reinforced composite structures are widely used in the aerospace, aircraft, civil and automotive applications due to their high strength-to-weight and stiffness-to-weight ratios and these applications require joining composite either to composite or to metal. There are three main methods for joining composite structures namely, bonding, mechanically fastened or a combination of the two. Bolted joint are preferred in structures where the disassembly is required for the purpose of maintenance and repair. Due to the stress concentration around the holes, bolted joints often represents the weakest part in the structure, and therefore it is important to design them safely. A review on the study of bolted joints in fiber reinforced composite structure is presented. It was found that the behavior of bolted joints in composite structure is affected by many factors, such as geometry, joint material, clamping–load provided by the bolts, ply orientations, etc. Accordingly, various researches have been conducted on the analyses of stress distribution, failure prediction, and strength properties of bolted joint both experimentally and numerically. Accurate prediction of stresses in bolted joints is essential for reliable design of the whole structure; if it is not optimally designed, premature and unexpected failures may be occurred.
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Hart-Smith, L. J. "An Engineer'S Viewpoint on Design and Analysis of Aircraft Structural Joints." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 209, no. 2 (April 1995): 105–29. http://dx.doi.org/10.1243/pime_proc_1995_209_278_02.
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The design of structurally efficient joints in aircraft fuselage structures and wing skin splices is addressed. It is contended that the joints should be designed first and the gaps in between filled in afterwards, taking pains not to optimize the basic structure first and then discover that it either cannot be assembled or that, when it is assembled, it is full of weak-link fuses. Both adhesively bonded and mechanically fastened joints are covered. Analogies are drawn between the characteristics of both classes of joints. The aspects of static joint strength and fatigue lives are included. The work is applicable to metallic as well as composite structures, and covers both high-load wing joints which have already been tested and new ideas for fuselage splices which have not. The effects of flaws and defects are associated with the need for damage tolerance, particularly in fuselage structures.
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Bautista Villamil, Alirio Andres, Juan Pablo Casas-Rodriguez, Alicia Porras Holguin, and Maribel Silva Barrera. "Mode I Crack Propagation Experimental Analysis of Adhesive Bonded Joints Comprising Glass Fibre Composite Material under Impact and Constant Amplitude Fatigue Loading." Materials 14, no. 16 (August 5, 2021): 4380. http://dx.doi.org/10.3390/ma14164380.
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The T-90 Calima is a low-wing monoplane aircraft. Its structure is mainly composed of different components of composite materials, which are mainly bonded by using adhesive joints of different thicknesses. The T-90 Calima is a trainer aircraft; thus, adverse operating conditions such as hard landings, which cause impact loads, may affect the structural integrity of aircrafts. As a result, in this study, the mode I crack propagation rate of a typical adhesive joint of the aircraft is estimated under impact and constant amplitude fatigue loading. To this end, effects of adhesive thickness on the mechanical performance of the joint under quasistatic loading conditions, impact and constant amplitude fatigue in double cantilever beam (DCB) specimens are experimentally investigated. Cyclic impact is induced using a drop-weight impact testing machine to obtain the crack propagation rate (da/dN) as a function of the maximum strain energy release rate (GImax) diagram; likewise, this diagram is also obtained under constant amplitude fatigue, and both diagrams are compared to determine the effect of each type of loading on the structural integrity of the joint. Results reveal that the crack propagation rate under impact fatigue is three orders of magnitude greater than that under constant amplitude fatigue.
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Tuo, Hongliang, Zhixian Lu, Xiaoping Ma, and Hongyu Guo. "Study on delamination damage evolution of composite L-shaped adhesive joint based on cohesive behavior." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 39, no. 2 (April 2021): 309–16. http://dx.doi.org/10.1051/jnwpu/20213920309.
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The adhesive joint of composite materials is one of the typical structures in aircraft structures, and the delamination damage is one of the most important damage modes in composite adhesive joints. In this paper, static and fatigue tests were carried out on L-shaped adhesive joints to analyze the damage evolution and failure modes of delamination damage under static and fatigue loadings. Based on the cohesive constitutive models, the static and high-cycle fatigue delamination constitutive models were developed. The static and fatigue numerical models of composite L-shaped adhesive joints were established by using finite element software. The stress distribution, deformation modes and delamination propagation laws were systematically studied. The simulation results are in good agreement with the experimental results. The delamination damage initiation, evolution and failure mechanism of L-shaped adhesive joints under static and fatigue loads were revealed by combining the experimental and the numerical results, which will provide theoretical and engineering guidance for strength and fatigue analysis of composite adhesive structures.
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Strizhius, V. E. "Fatigue life predictions of carbon fiber reinforced plastic in specimens of double-shear bolted joint." Civil Aviation High Technologies 24, no. 6 (December 27, 2021): 66–81. http://dx.doi.org/10.26467/2079-0619-2021-24-6-66-81.
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It is noted that in modern aircraft composite structures there is a significant number of composite and metal-composite shear bolted joints, the fatigue life of which is an important factor to ensure the operating safety of such constructions. Thus, special attention is given to the evaluation of the layered composites fatigue life in such joints during tests and calculations of the similar structures components. Despite a considerable number of publications and studies on this subject, it can be observed that many important methodological issues have not been solved yet in this field. These problems can deal with the choice of the main mode of layered composites fatigue damage in shear bolted joints; the uncertainty of the basic fatigue curve; the practical absence of some models, representing diagrams of constant life fatigue for the layered composites in the joints under consideration; the uncertainty of fatigue damage summation rule in the layered composites in the investigated joints. Based on the review results and the data analysis of domestic and foreign publications including the results of specially conducted studies, the solutions to these problems are proposed. The proposed solutions were verified by analyzing the calculated and experimental data on the fatigue life of carbon fiber reinforced plastic laminates НТА7/6376 [45/-45/0/90]3S in the double-shear bolted joints specimens.
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Tsokanas, Panayiotis, Theodoros Loutas, Dimitrios Pegkos, George Sotiriadis, and Vassilis Kostopoulos. "Mode II fracture toughness of asymmetric metal-composite adhesive joints." MATEC Web of Conferences 304 (2019): 01004. http://dx.doi.org/10.1051/matecconf/201930401004.
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The paper presents an experimental investigation of the mode II fracture toughness behavior of dissimilar metal-composite adhesive joints using the end-notched flexure (ENF) test. The adhesive joint under study consists of a thin titanium sheet joined with a thin CFRP laminate and is envisioned tobe applied in the hybrid laminar flow control system of future aircraft. Four different industrial technologies for the manufacturing of the joint areevaluated; co-bonding with and without adhesive and secondary bonding using either a thermoset or a thermoplastic composite. The vacuum-assisted resin transfer molding (VARTM) technique is employed for the manufacturing of the titanium-CFRP joint. After manufacturing, the joint is stiffened from its both sides with two aluminum backing beams to prevent large deformations during the subsequent ENF tests. Towards the fracture toughness determination from the experimental data, an analytical model recently reported by the authors is applied; that model considers the bending-extension coupling of each sub-laminate of the joint as well as the effect of the manufacturing-induced residual thermal stresses. The load-displacement behaviors, failure patterns, and fracture toughness performances for each of the four manufacturing options (MO) investigated are presented and compared.
15
Gao, Guoqiang, Luling An, Ioannis K. Giannopoulos, Ning Han, Ende Ge, and Geng Hu. "Progressive Damage Numerical Modelling and Simulation of Aircraft Composite Bolted Joints Bearing Response." Materials 13, no. 24 (December 8, 2020): 5606. http://dx.doi.org/10.3390/ma13245606.
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Finite element numerical progressive damage modelling and simulations applied to the strength prediction of airframe bolted joints on composite laminates can lead to shorter and more efficient product cycles in terms of design, analysis and certification, while benefiting the economic manufacturing of composite structures. In the study herein, experimental bolted joint bearing tests were carried out to study the strength and failure modes of fastened composite plates under static tensile loads. The experimental results were subsequently benchmarked against various progressive damage numerical modelling simulations where the effects of different failure criteria, damage variables and subroutines were considered. Evidence was produced that indicated that both the accuracy of the simulation results and the speed of calculation were affected by the choice of user input and numerical scheme.
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Pitta, Roure, Crespo, and Rojas. "An Experimental and Numerical Study of Repairs on Composite Substrates with Composite and Aluminum Doublers Using Riveted, Bonded, and Hybrid Joints." Materials 12, no. 18 (September 14, 2019): 2978. http://dx.doi.org/10.3390/ma12182978.
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In this work, experimental and numerical analyses of repairs on carbon fiber reinforced epoxy (CFRE) substrates, with CFRE and aluminum alloy doublers typical of aircraft structures, are presented. The substrates have a bridge gap of 12.7 mm (simulated crack), repaired with twin doublers joined with riveted, adhesive bonded, and hybrid joints. The performance of the repairs using different doubler materials and joining techniques are compared under static loading. The experimental results show that riveted joints have the lowest strength, while adhesive bonded joints have the highest strength, irrespective of the doubler material. Finite element analysis (FEA) of the studied joints is also performed using commercial FEA tool Abaqus. In the FEA model, point-based fasteners are used for the rivets, and a cohesive zone contact model is used to simulate the adhesive bond. The FEA results indicate that the riveted joints have higher tensile stresses on the metal doublers compared to the composite doublers. As per the failure modes, interestingly, for hybrid joints using composite doublers, the doublers fail due to net-section failure, while, for hybrid joints using metal doublers, it is the composite substrate that fails due to net-section failure. This suggests vulnerability of the composite structures to mechanical fastener holes. Lastly, the Autodesk Helius composite tool is used for prediction of first-ply failure and ply load distribution, and for progressive failure analysis of the composite substrate.
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Guo Shen, Wei, Xin Tong Zhao, and JunWei Han. "Composite Bonded Joints\' Lifetime for Aircraft under Random Fatigue Loads." Research Journal of Applied Sciences, Engineering and Technology 7, no. 15 (April 19, 2014): 3107–13. http://dx.doi.org/10.19026/rjaset.7.649.
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Ireman, Tomas, Tonny Nyman, and Kurt Hellbom. "On design methods for bolted joints in composite aircraft structures." Composite Structures 25, no. 1-4 (January 1993): 567–78. http://dx.doi.org/10.1016/0263-8223(93)90205-5.
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Liu, Longquan. "A study of the damage tolerance of composite-metal hybrid joints reinforced by multiple and penetrative thin pins." Composites and Advanced Materials 31 (January 2022): 263498332211055. http://dx.doi.org/10.1177/26349833221105523.
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The application of adhesive bonding technology in aircraft structures can reduce the total wight greatly, but the bonded joints are very sensitive to the possible manufacturing defects and damages during service operations, which makes them difficult to meet the damage tolerance requirements of the current transport airplane structures. In this study, the damage tolerance of composite-metal hybrid joints reinforced by multiple and penetrative thin pins was studied. The damage tolerance performance of the composite-metal joint is supposed to be enhanced by multiple through-the-thickness penetrative thin reinforcements in the bonding region, and the thin reinforcements were bonded together with both the composite and metallic joint plates. Both experimental tests and finite element simulations were conducted to investigate the effects of the through-the-thickness reinforcements on the damage tolerance performance of the joints with and without pre-fabricated disbond defects. Through the comparative analyses, it was found that the penetrative thin pins in the bonding region significantly improved the static load carrying capacity, the failure strain, the fracture energy, and the fatigue lives of the composite-metal bonded joints. Moreover, the reinforcements decreased the sensitivity of the bonded joints to the disbond defects in the bonding region. The damage tolerance performance of the composite-metal adhesively bonded joints was significantly increased by the through-the-thickness penetrative reinforcements and the enhancement mechanism was revealed by the combined analysis of test results and simulation results.
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Zhao, Shuyuan, Qian Sun, Yumin Zhang, and Jin Jia. "Parametric Influences of Geometric Dimensions on High Temperature Mechanical Behaviors and Damage Mechanisms of Ceramic Matrix Composite and Superalloy Double Bolted Joints." International Journal of Aerospace Engineering 2022 (August 31, 2022): 1–16. http://dx.doi.org/10.1155/2022/7169123.
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Given multiple material performance advantages, ceramic matrix composite (CMC) material has become one of the most promising hot structural materials used for thermal protection system in hypersonic vehicles. Under harsh thermal exposure of vehicles in flight, the design of connection structure would be a critical issue in improving load-carrying efficiency and ensuring service safety of aircraft structures in service environments. However, little attention was paid on mechanical behavior and its factors affecting the mechanical property of CMC joining at elevated temperature. To address this concern, a 3D finite element model coupled with progressive damage analysis is carried out to predict high temperature tensile properties and failure behavior of single-lap, double-bolt CMC/superalloy joints assembled by two serial protruding-head bolts. In the implementation of progressive damage analysis of 2D plain-woven C/SiC composites, a user-defined subroutine UMAT including a nonlinear constitutive model, 3D Alvaro failure criterion and Tan’s material degradation rule were embedded into the general package ABAQUS® through Fortran program interface. A parametric study considering geometries of joints was performed to evaluate their resultant influence on high temperature tensile behavior and the associated damage mechanisms for the CMC/superalloy double-bolt joint. New findings were provided for full exploitation of high performance through geometric design of ceramic matrix composite hot structure for hypersonic aircraft.
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Kim, Won Seok, and Jung Ju Lee. "Fracture Mechanics Characterization of Composite/Metal Interfaces of Bonded Joints." Key Engineering Materials 334-335 (March 2007): 361–64. http://dx.doi.org/10.4028/www.scientific.net/kem.334-335.361.
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Adhesive bonding between different materials has been widely used for a large variety of applications, such as in the aircraft, automotive, and many other civil engineering structures. Adhesive-bonded joints as load bearing components have the potential to save significant weight and cost over conventional riveted or bolted joints. For the last ten years a major problem in adhesive technology has been the difficulty in predicting the accurate load bearing capacity of a joint. This difficulty comes from the fact that the stress distribution in the adhesive joint is very complex and singular stress field exists at the bi-material corner. And for bonded joints, the failure usually occurs at the adhesive/adherend interface. Therefore another difficulty comes from the complex interfacial failure analysis due to the formation of chemical bonds, whose strengths are difficult to measure. Many studies have been conducted to investigate the effects of bond thickness, material properties of adhesives and adherends, and geometric shape of bi-material corner tip to the fracture behavior of bonded joints. In this paper, we analyze the stress fields at the interface corner of composite/steel(anisotropic/isotropic) double lap joint to predict failure by using stress intensity based fracture criterion. And analytical results are compared with experimental results of co-cured lap joints under tensile load condition. Micro-structural features, hardness characteristics, and fracture toughness determinations of the interfaces are also conducted.
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Álvarez, I., F. J. Doblas, C. Vallellano, A. Portal, and P. J. Arroyo. "FEA of the Influence of Assembly Parameters on the Fatigue Life of Metal-Composite Bolted Joints." Materials Science Forum 713 (February 2012): 73–78. http://dx.doi.org/10.4028/www.scientific.net/msf.713.73.
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This paper presents a numerical analysis of the influence on the fatigue life of hybrid metal/CFRP bolted joints of some deviations in certain assembly parameters. Two types of hybrid joints typically used in aircraft are analyzed, the Ti-6Al-4V/CFRP and the Al-2024-T3/CFRP, riveted with blind bolts. The parameters analyzed are: the thickness of sealant between metal sheet and composite laminate, the angle of countersunk and the adjustment or chamfer at the countersunk-drill zone. The analysis highlights critical areas for fatigue failure and the relative influence of these parameters on the fatigue resistance of the joint.
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Kostroun, Tomáš, and Milan Dvořák. "Application of the Pulse Infrared Thermography Method for Nondestructive Evaluation of Composite Aircraft Adhesive Joints." Materials 14, no. 3 (January 22, 2021): 533. http://dx.doi.org/10.3390/ma14030533.
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In this article, we examine the possibility of using active infrared thermography as a nontraditional, nondestructive evaluation method (NDE) for the testing of adhesive joints. Attention was focused on the load-bearing wing structure and related structural joints, specifically the adhesive joints of the wing spar caps and the skins on the wing demonstrator of a small sport aircraft made mainly of a carbon composite. The Pulse Thermography (PT) method, using flash lamps for optical excitation, was tested. The Modified Differential Absolute Contrast (MDAC) method was used to process the measured data to reduce the effect of the heat source’s inhomogeneity and surface emissivity. This method demonstrated a very high ability to detect defects in the adhesive joints. The achieved results are easy to interpret and use for both qualitative and quantitative evaluation of the adhesive joints of thin composite parts.
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Borba, N. Z., J. F. dos Santos, and S. T. Amancio-Filho. "Hydrothermal aging of friction riveted thermoplastic composite joints for aircraft applications." Composite Structures 255 (January 2021): 112871. http://dx.doi.org/10.1016/j.compstruct.2020.112871.
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Cheng, Liang, Qing Wang, and Yinglin Ke. "Experimental and numerical analyses of the shimming effect on bolted joints with nonuniform gaps." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 233, no. 11 (October 30, 2018): 3964–75. http://dx.doi.org/10.1177/0954406218809139.
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In order to investigate the effect of shim compensation for nonuniform gaps in aircraft assembly, the influence of the shims with different material and parameters on bolted joints is studied in this paper. According to the real material and assembly conditions of the aircraft joint structures, the specimen and experiment are designed to obtain the tensile performance of the joint structures with different shims. A three-dimensional finite element model, which incorporates the Johnson–Cook material property of the alloys, traction-separation law of liquid shims, contact relationships between the joint elements, and boundary conditions of the tensile process, is established with the specimen configurations. After validating through comparing with the experimental results, the modeling method is adopted to simulate the tensile response of the bolted joints with shims. Furthermore, both the influence of the shim material and thickness on the mechanical behaviors of bolted joints is investigated in detail. Shims can considerably reduce the assembly stress of joint structures and improve the joint stiffness and load capacity, and this effect is more remarkable with the increase of gap values. Liquid shims improve the joint stiffness due to its cohesive ability, while solid shims improve the joint load capacity. Hybrid shims possess a composite shimming effect of liquid and solid shims. Whatever the shim material is applied, the joint stiffness and strength drop with the growth of shim thickness, so strict deviation control method should be taken to ensure the assembly gaps as small as possible. The research results enhance the knowledge of shimming effect on joint structures, and thus offer positive guidance for practical application in aircraft assembly.
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Plaza, Borja, Daniel López, and David Poyatos. "Characterization of Joints between Carbon Fiber Composite Parts Using a Microstrip Transmission Line Method." Sensors 21, no. 4 (February 6, 2021): 1142. http://dx.doi.org/10.3390/s21041142.
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The electromagnetic performance of aerial platforms, which are composed mostly of nonmetallic materials, is a subject of great interest at present time. The behavior of this type of composite structure against electromagnetic environmental effects (E3), such as lightning, is not well-studied as in the case of metalic structures. The purpose of this article is to characterize the joints present in aerial platforms constructed mainly of nonmetallic composite materials. The study of these joints is fundamental because electrical discontinuities or preferential routes can produce changes in the electromagnetic behavior of an aircraft. The proposed measurement system for the characterization of these joints is a microstrip line. The flexibility of the test setup allows for evaluation of different joints in carbon fiber composite (CFC) samples with a different number of plies. Additionally, approximated models of the behavior of the joints as well as the detection of possible defects in the joining process are reported.
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McCarthy, M. A., V. P. Lawlor, and W. F. Stanley. "Effects of Variable Clearance in Multi-Bolt Composite Joints." Advanced Composites Letters 13, no. 4 (July 2004): 096369350401300. http://dx.doi.org/10.1177/096369350401300401.
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In this paper, the effects of variable clearance in multi-bolt, composite joints on load distribution, quasi-static strength and fatigue life have been studied experimentally. Both single-lap and double-lap configurations have been examined. The clearances examined ranged from neat-fit to clearances slightly larger than those allowed in the aircraft industry. Clearance has been found to have major effects on the load distribution in bolted joints, but negligible effect on ultimate strength. However, clearance had significant effects on initial failures such as bearing failure in one hole. A clearance-fit hole has been found to reduce fatigue life. The number of cycles to initiation of hole elongation under fatigue was particularly sensitive to the presence of a loose-fit hole.
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Kapidžić, Zlatan, Larsgunnar Nilsson, and Hans Ansell. "Finite element modeling of mechanically fastened composite-aluminum joints in aircraft structures." Composite Structures 109 (March 2014): 198–210. http://dx.doi.org/10.1016/j.compstruct.2013.10.056.
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Liu, Bin, Nongyue Gao, Bosen Tang, Yidi Gao, Peng Jin, Zhixian Lu, and Zaiguo Fu. "Progressive damage behaviors of the stepped-lap composite joints under bending load." Journal of Composite Materials 54, no. 14 (November 21, 2019): 1875–87. http://dx.doi.org/10.1177/0021998319889119.
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Due to the high connecting strength and excellent load-transferring efficiency, the adhesively stepped-lap composite bonding has been widely used as joints and repairs in advanced aircraft structures. In the service life of composite structures, they frequently subject to bending loads which are rarely reported by the available reference. Hence, in this paper, three-point and four-point bending experiments are performed on composite stepped-lap structures. Typical damage mode of adhesive on the stepped surface has been found in experiments. And the failure of composite fiber, matrix, interlamination and adhesive materials appeared as competitive behavior. Furthermore, numerical models of stepped-lap composite joints, which used cohesive zone model and 3D Hashin criteria to simulate the inter- and intralaminar damages, were utilized to predict the bending strength and progressive damage. The FEM model, based on progressive damage method, captured the detailed failure as the loading displacement was increasing. The simulation results of load displacement and damage have good agreement with the experiments.
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Yue, Xuande, Luling An, Zengtao Chen, Chufan Wang, Yuebo Cai, and Ruiheng Xiao. "Influence of gap filling on mechanical properties of composite-aluminum single-lap single-bolt hybrid joints." Advances in Mechanical Engineering 13, no. 2 (February 2021): 168781402198950. http://dx.doi.org/10.1177/1687814021989506.
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Composite-aluminum structures exist in aircraft structures generally. It is easy to cause gap between mating surfaces in composite-aluminum assembled structures with the curing deformation of composite. The composite-aluminum, single-lap, single-bolt joints were utilized to investigate the influence of forced assembly, liquid shim and peelable fiberglass shim on the mechanical properties of assembled structures. A steel gasket that removed the middle part was used in the joint to make a gap. The 3D Digital Imagine Correlation (3D-DIC) system was utilized to measure the strain field of specimens and the progressive damage model was created in ABAQUS. The results show that the shim filling can significantly increase the tension stiffness and peak load of the joint compare with forced assembly. As the shim thickness changes, the effects of the liquid shim and the peelable fiberglass shim on the tensile stiffness and peak load shows different. The liquid or peelable fiberglass shim can reduce the strain value around the hole and the peelable fiberglass shim has a better result than liquid shim. The squeeze between the bolt and composite laminate has a greater impact on matrix damage and fiber-matrix shear damage, while the secondary bending has a greater impact on matrix damage and fiber damage.
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Gröger, Benjamin, Daniel Köhler, Julian Vorderbrüggen, Juliane Troschitz, Robert Kupfer, Gerson Meschut, and Maik Gude. "Computed tomography investigation of the material structure in clinch joints in aluminium fibre-reinforced thermoplastic sheets." Production Engineering 16, no. 2-3 (December 7, 2021): 203–12. http://dx.doi.org/10.1007/s11740-021-01091-x.
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AbstractRecent developments in automotive and aircraft industry towards a multi-material design pose challenges for modern joining technologies due to different mechanical properties and material compositions of various materials such as composites and metals. Therefore, mechanical joining technologies like clinching are in the focus of current research activities. For multi-material joints of metals and thermoplastic composites thermally assisted clinching processes with advanced tool concepts are well developed. The material-specific properties of fibre-reinforced thermoplastics have a significant influence on the joining process and the resulting material structure in the joining zone. For this reason, it is important to investigate these influences in detail and to understand the phenomena occurring during the joining process. Additionally, this provides the basis for a validation of a numerical simulation of such joining processes. In this paper, the material structure in a joint resulting from a thermally assisted clinching process is investigated. The joining partners are an aluminium sheet and a thermoplastic composite (organo sheet). Using computed tomography enables a three-dimensional investigation that allows a detailed analysis of the phenomena in different joining stages and in the material structure of the finished joint. Consequently, this study provides a more detailed understanding of the material behavior of thermoplastic composites during thermally assisted clinching.
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Li, Binkai, Yu Gong, Hao Xiao, Yukui Gao, and Enquan Liang. "A Two-Dimensional Model for Pin-Load Distribution and Failure Analysis of Composite Bolted Joints." Materials 14, no. 13 (June 30, 2021): 3646. http://dx.doi.org/10.3390/ma14133646.
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Multibolt composite joints are widely used in aircraft structures. The determination of the pin-load distribution among the bolts is a critical step in the failure prediction of bolted joints. In this paper, a two-dimensional model of the multibolt composite structure is established for the pin-load distribution analysis. Its accuracy is validated by experimental results and the results from a 3D finite element model. Based on the determined pin-load distribution, FE models for a laminated plate with three-row fastener joints are built for failure prediction. Hashin stress criteria and the degradation guidelines of the material stiffness with respect to the different failure modes proposed by Tserpes are applied for the failure evaluation and the material degradation, respectively. The failure location and ultimate load are well predicted, which further validates the effectiveness and applicability of the proposed model for the pin-load distribution analysis.
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Sałaciński, Michał, Michał Stefaniuk, Piotr Synaszko, and Janusz Lisiecki. "Use of a composite repair patches to repair the upper air intake flap for the MiG-29 aircraft engine." Aviation Advances & Maintenance 40, no. 2 (December 1, 2017): 101–27. http://dx.doi.org/10.1515/afit-2017-0010.
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AbstractThe Composite Patch Bonded Repair (CPBR) is one of the most cost-efficient types of aircraft structure repair. In the CPBR, the damaged structure is reinforced by applying a composite patch. The boron-epoxy composite (BFRP) is a popular choice for these types of repair. The BFRP was utilized to repair the cracked resistance welding joints in the upper flap of the MiG-29’s RD-33 engine intake. In the present paper, the numerical results are shown, along with the comparative tests of the undamaged inlet flap, the damaged inlet flap and of the CPBR repaired inlet flap.
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Kashaev, Nikolai, Stefan Riekehr, Kay Erdmann, Alexandre Amorim Carvalho, Maxim Nurgaliev, Nikolaos Alexopoulos, and Alexandra Karanika. "Fracture mechanical behaviour of laser beam-welded AA2198 butt joints and integral structures." International Journal of Structural Integrity 6, no. 6 (December 7, 2015): 787–98. http://dx.doi.org/10.1108/ijsi-10-2014-0052.
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Purpose – Composite materials and metallic structures already compete for the next generation of single-aisle aircraft. Despite the good mechanical properties of composite materials metallic structures offer challenging properties and high cost effectiveness via the automation in manufacturing, especially when metallic structures will be welded. In this domain, metallic aircraft structures will require weight savings of approximately 20 per cent to increase the efficiency and reduce the CO2 emission by the same amount. Laser beam welding of high-strength Al-Li alloy AA2198 represents a promising method of providing a breakthrough response to the challenges of lightweight design in aircraft applications. The key factor for the application of laser-welded AA2198 structures is the availability of reliable data for the assessment of their damage tolerance behaviour. The paper aims to discuss these issues. Design/methodology/approach – In the presented research, the mechanical properties concerning the quasi-static tensile and fracture toughness (R-curve) of laser beam-welded AA2198 butt joints are investigated. In the next step, a systematic analysis to clarify the deformation and fracture behaviour of the laser beam-welded AA2198 four-stringer panels is conducted. Findings – AA2198 offers better resistance against fracture than the well-known AA2024 alloy. It is possible to weld AA2198 with good results, and the welds also exhibit a higher fracture resistance than AA2024 base material (BM). Welded AA2198 four-stringer panels exhibit a residual strength behaviour superior to that of the flat BM panel. Originality/value – The present study is undertaken on the third-generation airframe-quality Al-Li alloy AA2198 with the main emphasis to investigate the mechanical fracture behaviour of AA2198 BMs, laser beam-welded joints and laser beam-welded integral structures. Studies investigating the damage tolerance of welded integral structures of Al-Li alloys are scarce.
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Saravanan, Vimala, M. Ramachandran, and Chandrasekar Raja. "A Study on Aircraft Structure and Application of Static Force." 1 1, no. 1 (September 1, 2022): 1–6. http://dx.doi.org/10.46632/jame/1/1/1.
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In particular, the neural network approach to Demister-Schaefer theory has been demonstrated. Different modulus adherents, so that the adhesive pressures are approximately the same part of the vertical stabilizer was selected and reconstructed using the structure design of the hybrid wing body plane during the preliminary design phase. Due to the high number of design variables and various repairs, adhesive bonded fiber-composite joints or reinforcements are a challenging optimization problem, and this sheet directly affects structurally efficient and minimal damage by lightning strikes on aircraft structures. Parent systems rather than standard repairs. In the first part, the phenomenon of lightning curve connection in aircraft was introduced. Certain bonded joints offer more effective repairs than conventional mechanical joints, which are not fully credited with reducing their performance, and have important needs in most countries of the world. This article focuses on introducing integrated structural features and background, including application for civil aviation and flight qualification requirements, and the main purpose of this study is a computer-assisted materials selection tool. - Reinforced alloy. This paper explores the distribution of dynamic load identification given the unknown but limited uncertainties in aircraft configuration. This is the second of two documents on the structure of the Northwest Pacific Hurricane This sheet describes various structures Thin metal sheets are widely used to create different types of space structures due to its flexibility and ability to easily create any type of structure. Resource based to identify damage this article discusses the use of taxonomies.
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Tserpes, Konstantinos. "Numerical evaluation of crack stopping mechanisms in composite bonded joints due to corrugation and bolts." MATEC Web of Conferences 304 (2019): 01003. http://dx.doi.org/10.1051/matecconf/201930401003.
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In this paper, the crack stopping mechanisms in corrugated composite bonded joints and hybrid bonded/bolted joints were evaluated numerically using the cohesive zone modeling approach. For the study, the DCB (double-cantilever beam) and the CLS (crack-lap shear) specimens were modelled. The first two specimens were subjected to static loads and the latter both to static and fatigue loads. The analysis was performed using the LS-DYNA explicit FE code. Fatigue crack growth simulation was performed using an in-house developed user-defined subroutine (UMAT). The numerical results reveal a crack stopping in the corrugated DCB, no crack stopping in the corrugated CLS and a reduction of crack growth rate in the bonded/bolted CLS for both static and fatigue loads. The methods and the findings of the present study can be used for the design of crack stopping features in adhesively bonded primary composite aircraft structures.
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Stamatelos, Dimitrios G. "A Numerical Approach for Developing a Bearing-Bypass Design Criterion for Sizing Bolted Joints." Computation 10, no. 2 (January 20, 2022): 16. http://dx.doi.org/10.3390/computation10020016.
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Bolted joints are widely used in composite aircraft structures, for their assembly. The appropriate bolted joint configuration (hole/bolt diameter, pitch, etc.) is carefully selected during the detail design phase, where high fidelity numerical models are required with substantial computational cost and time. This work presents a design criterion, which allows the selection of the bolted joint configuration during the preliminary design phase with less computational time. The developed design criterion is based on a fully parametric finite element (FE) model, built in ANSYS V19 (Canonsburg, PA, USA), of a bolted joint with progressive damage modelling (PDM) capabilities, so that the failure of the joint can be predicted. From the numerical analyses, the bearing load and the load that bypasses the hole are calculated, up to failure, for a variety of joint configurations and loading conditions. The results of each analysis are used for plotting the failure envelope for the investigated bolted-joint configuration. Consequently, a design criterion is generated for the bolted joint. The availability of these failure envelopes, as design criterion, permit the appropriate selection of the bolted-joint configuration in an earlier design phase saving valuable time and computational cost.
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Liu, Hua, Weidong Zhu, Huiyue Dong, and Yinglin Ke. "An adaptive ball-head positioning visual servoing method for aircraft digital assembly." Assembly Automation 39, no. 2 (April 1, 2019): 287–96. http://dx.doi.org/10.1108/aa-05-2018-066.
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PurposeTo gain accurate support for large aircraft structures by ball joints in aircraft digital assembly, this paper aims to propose a novel approach based on visual servoing such that the positioner’s ball-socket can automatically and adaptively approach the ball-head fixed on the aircraft structures.Design/methodology/approachImage moments of circular marker labeled on the ball-head are selected as visual features to control the three translational degrees of freedom (DOFs) of the positioner, where the composite Jacobian matrix is full rank. Kalman–Bucy filter is adopted for its online estimation, which makes the control scheme more flexible without system calibration. A combination of proportional control with sliding mode control is proposed to improve the system stability and compensate uncertainties of the system.FindingsThe ball-socket can accurately and smoothly reach its desired position in a finite time (50s). Positional deviations between the spherical centers of ball-head and ball-socket in theX-Yplane can be controlled within 0.05 mm which meets the design requirement.Practical implicationsThe proposed approach has been integrated into the pose alignment system. It has shown great potential to be widely applied in the leading support for large aircraft structures in aircraft digital assembly.Originality/valueAn adaptive approach for accurate support of large aircraft structures is proposed, which possesses characteristics of high precision, high efficiency and excellent stability.
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Топал, Микола Савович, and Володимир Михайлович Андрющенко. "ПОШКОДЖЕННЯ ВІД ЩІЛИННОЇ КОРОЗІЇ В КОНСТРУКЦІЯХ ЛІТАКІВ ТА ЇХ ВИЯВЛЕННЯ." Open Information and Computer Integrated Technologies, no. 83 (May 23, 2019): 143–55. http://dx.doi.org/10.32620/oikit.2019.83.10.
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Presented examples of destruction of aircraft designs due to corrosion of metals under conditions of fatigue loading. It is shown that slit corrosion, which is an increase in corrosion in crevice and gaps between two metals, as well as in places of untight contact of metal with a nonmetallic material resistant to corrosion, leads to the appearance of corrosion products in the joints of the skin with the power suite, which supports it , which can lead to the swelling of some elements of the joint relative to other elements and provoke the tearing off of the heads of rivets with the further development of fatigue cracks and the destruction of aircraft structures. Shown, that visual inspection is not always effective for the detection of corrosion damage, and sometimes impossible, for example, in closed internal structures. New developments in the field of sensors and equipment for the detection of corrosive substances and corrosion damage are presented. Among them is information on the sensor (organic-ceramic composite) containing the conducting complex. When the composite is exposed to water liquids, its conductivity is lost. When the composite dries, the sensor reaches its initial values of resistance. Information is provided on the optical sensor for detecting corrosion in the construction of the airframe. This sensor is based on the remote detection of aluminum ions formed during corrosion. The development of a multi-parameter integrated sensor for assessing the structural integrity of aluminum alloys, the recording of the concentration of chloride ions, the release of hydrogen, changes in humidity and degradation of the material is presented. Information is provided on fluorescence-based optical sensors used to detect specific ions such as aluminum, indicating the beginning of corrosion of an aluminum alloy. Information is provided on the development of advanced digital X-ray methods for the detection of corrosion in the design of aircraft. The conclusion is made on the necessity of combining visual control and control with the use of means and methods for detecting corrosive substances and corrosion damage.
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Мальцев, О. Ю., С. А. Толстой, Д. І. Конотоп, and В. В. Сухов. "ОСОБЛИВОСТІ СКЛАДАННЯ ПЛАНЕРІВ ЛІТАКІВ, ЯКІ МІСТЯТЬ ЕЛЕМЕНТИ ІЗ КОМПОЗИЦІЙНИХ МАТЕ-РІАЛІВ." Open Information and Computer Integrated Technologies, no. 95 (October 26, 2022): 29–44. http://dx.doi.org/10.32620/oikit.2022.95.03.
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The results of the analysis of the possibility of using known methods for assembling the components of an aircraft airframe, containing, in particular, elements of composite materials (CM), are presented. The features of structures made of CM and their joints, which affect the subsequent assembly of such structures, are considered. Systematized information about the terminology in the home area. Alternative formulations of assembly-related terms are proposed.Some features of the choice of methods for assembling structures containing elements from CM are considered. The properties of CM that influence the choice of assembly bases are analyzed. Potentially possible bases for assembling components of aircraft structures from CM, as well as CM and metals in different proportions, in particular, the method of assembly using virtual bases, are proposed. The classification of methods for assembling aircraft all-metal structures is carried out, technological limitations of CM are indicated, which have a direct impact on the use of the corresponding elements of assembly equipment. Some cases of imitation of design solutions of aircraft structures made of metals into structures containing elements of CM are considered. The typification of the main assembly units of the airframe of the aircraft has been made, in terms of orientation to certain assembly methods.Formulas for calculating the value of the total error in the assembly of components made of metals are proposed. The factors influencing the magnitude of the total error for CM structures are substantiated. The current trends in the use of CM at the SE «ANTONOV» are formulated in an enlarged way. An adaptive scheme for choosing a method for assembling assembly units containing, in particular, elements from CM, obtained by extrapolation from all-metal structures, has been developed.Conclusions are drawn about the need to orient the developed aircraft structure to a certain assembly method, and the main conditions for carrying out assembly work on composite parts from CM are formulated. The possibility of assembling aircraft structures from CM, as well as CM and metals in different proportions, with predominantly based on holes using simple assembly equipment, has been determined.
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Arena, Maurizio, Antonio Chiariello, Martina Castaldo, and Luigi Di Palma. "Vibration Response Aspects of a Main Landing Gear Composite Door Designed for High-Speed Rotorcraft." Aerospace 8, no. 2 (February 19, 2021): 52. http://dx.doi.org/10.3390/aerospace8020052.
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One of the crucial issues affecting the structural safety of propeller vehicles is the propeller tonal excitation and related vibrations. Propeller rotation during flight generates vibrating sources depending upon its rotational angular velocity, number of blades, power at shaft generating aircraft thrust, and blade geometry. Generally, the higher energy levels generated are confined to 1st blade passing frequency (BPF) and its harmonics, while additional broadband components, mainly linked with the blade shape, the developed engine power, and the turbulent boundary layer (TBL), also contribute to the excitation levels. The vibrations problem takes on particular relevance in the case of composite structures. The laminates in fact could exert damping levels generally lower than metallic structures, where the greater amount of bolted joints allow for dissipating more vibration energy. The prediction and reduction of aircraft vibration levels are therefore significant considerations for conventional propeller aircrafts now entering the commercial market as well as for models currently being developed. In the Clean Sky 2 framework, the present study focuses on a practical case inherent to the AIRBUS-Racer program aiming to design and develop a multi-tasking fast rotorcraft. This paper defines a finite elements (FE)-based procedure for the characterization of the vibration levels of a main landing gear (MLG) composite door with respect to the expected operating tonal loads. A parametric assessment was carried out to evaluate the principal modal parameters (transfer functions and respective resonance frequencies, mode shapes, and damping coefficients) of the landing gear-door assembly in order to achieve reduced vibration levels. Based on the FE analysis results, the influence of the extra-damping, location, and number of ballast elements, the boundary conditions were investigated with respect to failure scenarios of the kinematic line opening the study towards aeroelastic evaluations. Further experimental ground test results serve as a validation database for the prediction numerical methods representative of the composite door dynamic response.
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Luan, Chao, Kai Fu Zhang, Xiao Ge Li, and Yuan Li. "A Method for Predicting the Onset and Expanding of the Damage during CFRP Interference-Fit Joining Process." Advanced Materials Research 567 (September 2012): 166–69. http://dx.doi.org/10.4028/www.scientific.net/amr.567.166.
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Interference-fit is the main joint mode of the aircraft composite components, however, improper interference-fit percentage will lead to the onset and the expanding of the damage around the hole. Currently, the interference-fit percentage is decided only through experience, and theoretical foundation is lacked. This paper first analyzes the stress condition of the interference-fit joining process, and constructed the stress distribution model based on interference-fit percentage. Based on this, damage onset model is constructed combined with the Yamada criterion, and the damage expanding model is constructed with the help of the Characteristic curve method; Secondly ABAQUS is used to simulate the interference-fit joining process; Finally contrast is made between the numerical and experimental of joints with different diameters, and the result of the contrast shows that the damage region onset and expanding model and the degeneration model are good.
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Devitte, Cristiano, Gabriel S. C. Souza, André J. Souza, and Volnei Tita. "Optimization for drilling process of metal-composite aeronautical structures." Science and Engineering of Composite Materials 28, no. 1 (January 1, 2021): 264–75. http://dx.doi.org/10.1515/secm-2021-0027.
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Abstract Metal-composite laminates and joints are applied in aircraft manufacturing and maintenance (repairing) using aluminum alloys (AA) and glass fiber-reinforced polymer (GFRP). In these applications, drilling has a prominent place due to its vast application in aeronautical structures’ mechanical joints. Thus, this study presents the influence of uncoated carbide drills (85C, 86C, H10N), cutting speeds (v c = 20, 40, and 60 m min−1), and feed rates (f = 0.05, 0.15, and 0.25 mm rev−1) on delamination factor, thrust force ( F t {F}_{\text{t}} ), and burr formation in dry drilling metal-composite laminates and joints (AA2024/GFRP/AA2024). Experiments were performed, analyzed, and optimized using the Box–Behnken statistical design. Microscopic digital images for delamination evaluation, piezoelectric dynamometer for thrust force acquisition, and burr analysis were considered. The major finding was that the thrust force during drilling depends significantly on the feed rate. Another significant factor was the influence of the drill type (combined or not with feed rate). In fact, it was verified that the feed rate and the drill type were the most significant parameters on the delamination factor, while the feed rate was the most relevant on thrust force. The cutting speed did not affect significantly thrust force and delamination factor at exit ( F da S ) \hspace{.25em}({F}_{{\text{da}}_{\text{S}}}) . However, the combination f × v c was significant in delamination factor at entrance ( F da E ) \text{ }({F}_{{\text{da}}_{\text{E}}}) . Based on the optimized input parameters, they presented lower values for delamination factors ( F da E = 1.18 {F}_{{\text{da}}_{\text{E}}}=1.18 and F da S = 1.33 {F}_{{\text{da}}_{\text{S}}}=\hspace{.25em}1.33 ) and thrust force ( F t = 67.3 N {F}_{\text{t}}=67.3\hspace{.5em}\text{N} ). These values were obtained by drilling the metal-composite laminates with 85C-tool, 0.05 mm rev−1 feed rate, and 20 m min−1 cutting speed. However, the burrs at the hole output of AA2024 were considered unsatisfactory for this specific condition, which implies additional investigation.
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Sagomonova, V. A., S. S. Dolgopolov, A. E. Sorokin, and V. V. Tselikin. "EVALUATION OF VIBRATION-ABSORBING MATERIAL BASED ON POLYURETHANE USING POSSIBILITY FOR A SEAL." Proceedings of VIAM, no. 10 (2021): 28–35. http://dx.doi.org/10.18577/2307-6046-2021-0-10-28-35.
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The article presents the results of a study of the operational properties of the VTP-1V grade material, including cold flow. It was found that the material of the VTP-1V grade hasn’t cold flow and, according to the complex of properties studied in the work, can be recommended as an alternative to layered polymeric composite materials used in the bolted joints of the aircraft wing rib. The possibility of operating VTP-1V in an environment of aggressive liquids was assessed and it was shown that the material is resistant to the effects of TC-1 fuel, including retaining low deformability under prolonged loading.
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Han, Qing, Bin Liu, and Wei Xie. "The tension failure mechanisms of composite stepped bonding repairs and joints for the aircraft primary load-bearing structures." Journal of Adhesion Science and Technology 33, no. 7 (January 20, 2019): 675–90. http://dx.doi.org/10.1080/01694243.2018.1558477.
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Baker, Wayne, Iain McKenzie, and Rhys Jones. "Development of life extension strategies for Australian military aircraft, using structural health monitoring of composite repairs and joints." Composite Structures 66, no. 1-4 (October 2004): 133–43. http://dx.doi.org/10.1016/j.compstruct.2004.04.031.
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OLIVEIRA, VITÓRIA SIMPLÍCIO, RAFAEL RESENDE LUCAS, TAYENNE PRADO CARVALHO, LUIS FELIPE MARQUES, JONAS FRANK REIS, ANA BEATRIZ RAMOS MOREIRA ABRAHÃO, and EDSON COCCHIERI BOTELHO. "Development of the Oxyacetylene Welding Process for PEI/Glass Fiber Laminates." Welding Journal 100, no. 4 (April 1, 2021): 142–49. http://dx.doi.org/10.29391/2021.100.012.
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The technology for joining thermoplastics through welding offers numerous advantages over mechanical joining. Currently, the joining of composite parts with weight reduction and cost savings is being developed to improve aircraft performance. This paper proposes the use of oxygen-acetylene as a process for bonding composite materials. Oxyacetylene welding is a simple and economical method that can be suitable for polymeric materials. The advantage of applying this technique is a more accessible process that is composed of a portable system with low cost. In evaluating the welding efficiency for composite materials, the lap shear strength (LSS) mechanical test stands out among the most referenced essays in the literature. This work aimed to study the development of oxyacetylene flame welding as well as the optimization of welding parameters for polyetherimide/glass fiber composite. The optimization was performed using complete factorial planning 22 as a tool, and the variables studied were time and distance of the flame. With the optimized condition set as the response variable with the highest lap shear value, the joints obtained were measured for their quality by means of end-notched flexure mechanical testing, thermal analysis, and fracture analysis after LSS testing using optical and electronic microscopy.
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Stamatelos, Dimitriοs, and George Labeas. "Towards the Design of a Multispar Composite Wing." Computation 8, no. 2 (April 9, 2020): 24. http://dx.doi.org/10.3390/computation8020024.
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In the pursuit of a lighter composite wing design, fast and effective methodologies for sizing and validating the wing members (e.g., spar, ribs, skins, etc.) are required. In the present paper, the preliminary design methodology of an airliner main composite wing, which has an innovative multispar configuration instead of the conventional two-spar design, is investigated. The investigated aircraft wing is a large-scale composite component, requiring an efficient analysis methodology; for this purpose, the initial wing sizing is mostly based on simplified Finite Element (FE) stress analysis combined to analytically formulated design criteria. The proposed methodology comprises three basic modules, namely, computational stress analysis of the wing structure, comparison of the stress–strain results to specific design allowable and a suitable resizing procedure, until all design requirements are satisfied. The design constraints include strain allowable for the entire wing structure, stability constraints for the upper skin and spar webs, as well as bearing bypass analysis of the riveted/bolted joints of the spar flanges/skins connection. A comparison between a conventional (2-spar) and an innovative 4-spar wing configuration is presented. It arises from the comparison between the conventional and the 4-spar wing arrangement, that under certain conditions the multispar configuration has significant advantages over the conventional design.
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Tserpes, Konstantinos, and Elli Moutsompegka. "The Effect of Pre-Bond Contamination by Thermal Degradation and De-Icing Fluid on the Tensile Strength of Scarf Composite Bonded Joints." Journal of Composites Science 5, no. 7 (June 28, 2021): 168. http://dx.doi.org/10.3390/jcs5070168.
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The synergistic effect of pre-bond contamination by thermal degradation and de-icing fluid on the tensile behavior of scarf composite bonded joints has been investigated experimentally. The contamination types considered are related to the repair process of composite aircraft structures. Three contamination scenarios have been considered: namely, thermal degradation (TD) and a combination of thermal degradation with two different levels of de-icing fluid (TD+DI1 and TD+DI2). DI2 is more severe than DI1. Contamination has been applied to one of the adherents while the other one has been intentionally left intact. Tension tests have been conducted on single-lap shear specimens. The experimental results were compared with the reference samples (REF) showing an increase in tensile strength for the TD specimens and a decrease in tensile strength for the TD+DI1 and TD+DI2 specimens. After the tension tests, the failure surfaces were evaluated to get a better insight of the failure mechanisms of the bondline and to assess the effect of contamination. The TD specimens presented an increased cohesive failure which is consistent with the increase of the failure load, while the combined contamination caused the failure of the composite adherents which again is consistent with the decrease of tensile strength of the scarf specimens.
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Kondapalli, Siva Prasad, Srinivasa Rao Chalamalasetti, and Nageswara Rao Damera. "Optimization of Fusion Zone Grain Size, Hardness, and Ultimate Tensile Strength of Pulsed Current Microplasma Arc Welded AISI 304L Sheets Using Genetic Algorithm." International Journal of Manufacturing Engineering 2014 (March 5, 2014): 1–8. http://dx.doi.org/10.1155/2014/943643.
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Austenitic stainless steel sheets have gathered wide acceptance in the fabrication of components, which require high temperature resistance and corrosion resistance, such as metal bellows used in expansion joints in aircraft, aerospace, and petroleum industry. In case of single pass welding of thinner sections of this alloy, Pulsed Current Microplasma Arc Welding (PCMPAW) was found beneficial due to its advantages over the conventional continuous current process. The quality of welded joint depends on the grain size, hardness, and ultimate tensile strength, which have to be properly controlled and optimized to ensure better economy and desirable mechanical characteristics of the weld. This paper highlights the development of empirical mathematical equations using multiple regression analysis, correlating various process parameters to grain size, and ultimate tensile strength in PCMPAW of AISI 304L sheets. The experiments were conducted based on a five-factor, five-level central composite rotatable design matrix. A genetic algorithm (GA) was developed to optimize the process parameters for achieving the desired grain size, hardness, and ultimate tensile strength.