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Dissertations / Theses on the topic 'Composite aircraft'

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Published: 4 June 2021
Last updated: 7 February 2022

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1

Ashawesh, Gamal Mohamed. "Flutter behaviour of composite aircraft wings." Thesis, Cranfield University, 1999. http://hdl.handle.net/1826/3900.

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This research work presents series of investigations into the structural dynamics and dynamic aeroelastic (flutter) behaviour of composite and metal wings. The study begins with a literature review where the development and an over view of the previous investigations in this field are presented. Static stiffness is very important to any type of analysis, especially in both dynamic and flutter analysis as in this case. Therefore, different methods are presented and used for the determination of cross- sectional rigidities such as bending, torsional and bending-torsional coupling rigidities properties for beams constructed of laminated and thin-walled structures materials. A free vibration experimental analysis was conducted on the physical Cranfield Al aerobatic composite wing box structure. The composite wing box was exited in the frequency range of 0 to 300 Hz, with both sinusoidal and random excitations, which yields to six resonant frequencies. The theoretical free vibration and flutter analysis was then carried out firstly on the physical Cranfield Al aerobatic metal wing box. The metal wing was modeled using two techniques; the first model was a simplified wing structure (beam with lumped mass). This analysis of the simplified model was done using CALFUN program for the free vibration analysis and using MSC/NASTRAN for both free vibration and flutter analysis. The second model was a detailed model created by MSC/PATRAN and analyzed by MSC/NASTRAN for the free vibration and flutter analysis. The obtained results (natural frequencies and mode shapes) showed a good agreement between the simplified, detailed model and the experimental test. It was found that even with using the simplified model, but having the physical characteristics of the wing leads to a good agreement with the detailed model and experimental work. This also showed the importance of simplified model at early stage of the design to the structural designer in terms of the accuracy, time, and size of the model. Free vibration and flutter analysis was carried out on the Cranfield Al aerobatic composite wing box with the original laminate lay ups using Lanczos method for extracting the eigenvalues and eigenvectors and using PK method for finding the flutter speed and frequency provided by MSC/NASTRAN. The results were compared with the experimental vibration analysis and were found a large difference in the first frequency mode. To investigate the cause of the variation, a series of static loading tests were performed on the composite wing box. Also a comparison of the results between the metal and composite aerobatic wing box is presented. It was found that the large difference could be due to the combination of different parameters such as stiffness (age of the wing, delamination and boundary condition), and increase of mass of the physical wing box (due to environmental effect such as moisture) and modelling differences. The free vibration characteristics of ten wing models constructed from balanced and unbalanced laminate configurations were carried out using Lanczos method provided by MSC/NASTRAN. The analysis was done on ten wing models modeled to simulate Circumferentially Asymmetric Stiffness (CAS) and Circumferentially Uniform Stiffness (CUS). The static equivalent stiffness was calculated using two different modeling methods for a wide range of fibre angles 0 (- 90° to 90°) of the skins. The variations and the importance of the stiffness ratio (EI/GJ), parameter (K/GJ), and the frequency ratio (wb/(Ot) are illustrated against the fibre angle 0. It was found that the fundamental bending frequency is slightly lower in the case of CUS (K = 0) as compared to the CAS (K # 0), which was not the case in the plate model. Also, the first torsion frequency mode in the case of CUS was much lower than the first torsion frequency of the CAS, which was not the case of the plate model. However, the effect of bend-twist coupling stiffness on the mode shapes was pronounced in both structures especially at the area of higher coupling stiffness. The flutter analysis was done using the PK method for all the wing models of both (CAS) and (CUS) configurations. The results showed the optimum value of flutter speed and the importance of the stiffness ratio (EI/GJ), parameter (K/GJ), the frequency ratio (wb/wt), which will lead to the maximum flutter speed. The effects of the above parameters, geometrical coupling and the wash-in and washout on the non- dimensional flutter speed are presented. It was concluded that, negative bend-twist coupling stiffness is beneficial for flutter speed compared to the positive bend-twist coupling stiffness at 00<0<_30°. It was also found that the flutter speed for the CUS was higher at 00<0<_300 compared to the CAS. Also creating an offsite between the elastic axis and center of gravity (behind) decreases the flutter speed whereas having more ribs increases the flutter speed compared with adding stringers. The analysis was carried out on a more practical composite wing box, which was the physical Cranfield Al aerobatic composite wing box. There are some simplifications on the physical structure, which are the cancellation of the woven materials and keeping the same laminate lay ups for the upper and lower skin. The natural frequency and mode shapes was obtained and plotted against the fibre angle 0 of the upper and lower skin for the (CAS) and (CUS) configurations using both symmetric and asymmetric laminate for the upper and lower skin. The flutter analysis was done for the composite wing box for the same configurations as in the free vibration analysis. The effects of the fibre angle 0 of the upper and lower skin, material coupling stiffness, wash-in and wash-out, and structural damping on the non- dimensional flutter speed and flutter frequencies are illustrated. It was found that in this configurations both structural and bend-twist coupling are exist, negative bend- twist coupling (wash-in) increases the flutter speed compared with the positive bend- twist coupling, and the possibility of increasing the flutter speed at higher frequency ratio, structural coupling and positive bend-twist coupling (wash-out).
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2

Taylor, James Marcus. "Optimisation and validation of frequency constrained composite wings." Thesis, University of Bath, 1998. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.268213.

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3

Stephens, V. M. "Crashworthiness of composite seats for civil aircraft." Thesis, Cranfield University, 1992. http://hdl.handle.net/1826/1771.

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A study has been conducted into the design of civil aircraft seats which are forward-facing and use the lap-belt method of restraint. Within these terms of reference, the response of the seat restraint occupant system (SROS) to impact loading has been analysed using physical (dynamic testing) and analytical (computer simulation) modelling techniques. With the increasing use of fibre-reinforced polymer composites in aircraft for weight efficiency, and the consequent appearance of composite seats, attention must be given to the crash performance of these structures. Composite structures are characterised by brittle failure with low impact energy absorption, in comparison to the collapse of metal structures which may exhibit plastic deformation prior to failure. However, using the developing technology of composite sub-structures with high specific energy absorption capability, seat structures have been modified to incorporate composite load-limiting elements. The redesign process involved the compatibility of energy absorber loads with occupant dynamics to minimise injury potential, together with the alleviation of forces in the structural load path to reduce damage and preclude failure of the seat, floor track, and other components. Shortcomings of existing seat designs were assessed, and the dynamics of lap-belted occupants analysed, including secondary head impact with the forward seat structure. The computer' model created was validated against the results of dynamic tests, and then used in a parametric study of occupant dynamics. Conclusions and recommendations include guidlines relating to the future design of both metal and composite seats.
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4

Cook, Lawrence. "Visual inspection reliability for composite aircraft structures." Thesis, Cranfield University, 2009. http://dspace.lib.cranfield.ac.uk/handle/1826/6834.

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This thesis presents a study of the effects of surface colour, surface finish and dent shape on the visual inspection reliability of 3D surface indentations common in shape to those produced by impact damage to carbon fibre reinforced epoxy laminates. Falling weight (2.5kg) apparatus was used to produce impact damage to non-painted, non-mesh Hexcel AS4/ 8552 carbon fibre reinforced plastic (CFRP) laminates and painted AS4/ 8552 laminates containing bronze mesh and glass fabric lightning strike protection layers. Ø20 mm and Ø87 mm hemispherical tip impacts to painted 17ply and 33ply laminates at varying energy levels typically produced circular shaped, smoothly contoured, rounded sectional profiles with an absence of surface breaking cracks. Sectional profiles through coordinate measuring (CMM) data of the impact dents were described using a set of geometric variables. Identifying relationships between impact energy and the geometric variables allowed the typical sectional profile through impact damage dents from Ø20 mm and Ø87 mm hemispherical tips on 17ply and 33ply painted CFRP laminates to be calculated for energies between 5J to 80J. Calculated sectional profiles typical of impact damage dents to CFRP laminates were reconstructed as simple revolved shapes using 3D computer aided design (CAD) models. The 3D CAD models were computer numerical control (CNC) machined into 3mm Plexiglas panels to produce facsimiles of hemispherical impact damage dents on CFRP laminates. Facsimile specimen sets of sixteen 600 mm x 600 mm panels were produced in gloss and matt grey, white and blue finishes. Each set contained the same 32 different sized machined dents representing Ø20 mm and Ø87 mm hemispherical tip impact damage to 17ply & 33ply painted CFRP laminate. Each facsimile specimen set was combined with similarly finished unflawed (dent free) panels. 64 panels in each colour/ finish were presented for 5 seconds in a randomised order to a minimum of 15 novice participants in a visual inspection task lasting approximately 25 minutes. II A set of corresponding visual inspection experiments were performed in which physical specimens were replaced by digitally projected actual size photorealistic images of the machining CAD data. Comparisons between the results of the physical and virtual specimen trials revealed differences in detectability for similarly sized dents. The detection results obtained from visual inspection of physical specimens demonstrated that the detectability of dents similar to those caused by higher (>40J) energy impacts from a Ø87 mm hemispherical tip was less than that of the dents caused by lower energy (<20J) impacts from Ø20 mm tips. However, larger subsurface delamination area was demonstrated by the higher energy Ø87 mm impacts than lower energy Ø20 mm impacts on 150 mm x 100 mm coupons of the same thickness laminate. The results of these experiments imply that detectability of dents caused by larger diameter objects at higher energies cannot be assumed to be greater than that of lower energy impacts from smaller diameter objects. The detection results demonstrate that detectability by visual inspection cannot be assumed the same for an impact dent on different surface colours and finishes. In general terms, the highest numbers of dents returning >90% detection were observed on grey specimens and the highest number of dents returning 0% detection were observed on matt blue specimens. The difference in detection rates for similarly sized dents on a gloss and matt finish was least on grey coloured specimens and greatest on blue coloured specimens.
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5

Lillico, Mark. "Aeroelastic optimisation of composite wings." Thesis, University of Bath, 1997. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.362287.

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6

Bigand, Audrey. "Damage assessment on aircraft composite structure due to lightning constraints." Thesis, Toulouse, ISAE, 2020. http://www.theses.fr/2020ESAE0027.

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L’utilisation des matériaux composites dans l’industrie aéronautique s’étant largement étendue, ledimensionnement de ces structures et de leur protection vis-à-vis de la foudre est devenu un enjeu majeur. Ilest important de pouvoir développer des outils prédictifs permettant d’obtenir une conception de structurerépondant aux critères de certification avec des temps et coûts de conception maitrisés. L’interaction de lafoudre avec une structure composite est un phénomène multiphysique complexe, avec une difficulté ajoutéepar la présence d’une protection métallique en surface et d’une couche de peinture. Dans ce contexte, cetteétude a visé à développer la compréhension par rapport aux forces générées par la foudre et d’en évaluer sesconséquences quant à l’endommagement du composite. Dans cet objectif, le phénomène a d’abord étédécomposé pour en étudier ses différentes parties et définir l’impact des interactions. Dans un premier temps,l’arc libre a été comparé au pied d’arc en interaction avec différents substrats permettant de définir un modèlede vaporisation de la protection foudre. Dans un second temps, la surpression générée par l’explosion de laprotection en surface lors de la vaporisation a été évaluée pour définir des profils de pression spatio-temporels.Dans un troisième temps, une caractérisation mécanique de la peinture a été développée afin de quantifier soneffet de confinement sur l’explosion de surface. A chaque étape, une théorie a été développée et analysée viades modèles numériques et des essais. Enfin, ces trois différentes briques ont été rassemblées dans un modèlemécanique simulant l’impact foudre sur une structure composite afin d’en prédire l’endommagement. De plus,une loi utilisateur a été développée pour appliquer ce chargement complexe ainsi qu’une loid’endommagement. Ces modèles sont comparés aux résultats d’essai foudre en laboratoire afin d’endéterminer les limites de validité et leur capacité à prédire l'endommagement<br>As composite materials are now widely used in the aeronautical industry, the sizing of these structures andtheir protection against lightning has become a major issue. It is important to develop predictive tools to obtaina structure concept that meets certification requirements with a controlled time and cost during the designphase. The interaction of lightning with a composite structure is a complex multi-physics phenomenon, with afurther difficulty due to the presence of a metallic protection on the surface and a layer of paint. In this context,this study aimed to develop an understanding of the forces generated by lightning and to assess itsconsequences in terms of damage to the composite. To this end, the phenomenon was first broken down tostudy its different components and define the impact of their interactions. In a first step, the free arc wascompared to the arc root in interaction with different substrates to define a vaporisation model of the lightningprotection. In a second step, the overpressure generated by the explosion of the surface protection duringvaporisation was evaluated to define spatio-temporal pressure profiles. In a third step, a mechanicalcharacterization of the paint was developed in order to quantify its confinement effect on the surface explosion.At each stage, a theory was developed and analysed via numerical models and tests. Finally, these threedifferent bricks are brought together in a mechanical model simulating the lightning impact on a compositestructure in order to predict the damage. In addition, a user subroutine has been developed to apply thiscomplex loading as well as a damage law. These models are compared with lightning laboratory test results todetermine their validity limits and their ability to predict the damage
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7

Nyman, Tonny. "Fatigue and residual strength of composite aircraft structures." Doctoral thesis, KTH, Aeronautical Engineering, 1999. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-2848.

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8

Whisler, Daniel A. "Low velocity blunt impacts on composite aircraft structures." Diss., [La Jolla] : University of California, San Diego, 2009. http://wwwlib.umi.com/cr/ucsd/fullcit?p1470612.

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Thesis (M.S.)--University of California, San Diego, 2009.<br>Title from first page of PDF file (viewed January 12, 2010). Available via ProQuest Digital Dissertations. Includes bibliographical references (p. 102-104).
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9

Ahmad, M. K. M. "Shear lag effect in composite box girders." Thesis, Cardiff University, 1989. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.237869.

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10

Ericsson, Max. "Simulating Bird Strike on Aircraft Composite Wing Leading Edge." Thesis, KTH, Hållfasthetslära (Inst.), 2012. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-103783.

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In this master thesis project the possibility to model the response of a wing when subjected to bird strike using finite elements is analyzed. Since this transient event lasts only a few milliseconds the used solution method is explicit time integration. The wing is manufactured using carbon fiber laminate. Carbon fiber laminates have orthotropic material properties with different stiffness in different directions. Accordingly, there are damage mechanisms not considered when using metal that have to be modeled when using composites. One of these damage mechanisms is delamination which occurs when cured layers inside a component become separated. To simulate this phenomenon, multiple layers of shell elements with contact in between are used as a representation of the interface where a component is likely to delaminate. By comparing experimental and simulated results the model of delamination is verified and the influence of different parameters on the results is investigated. Furthermore, studies show that modeling delamination layers in each possible layer of a composite stack is not optimal due to the fact that the global stiffness of the laminate is decreased as more layers are modeled. However, multiple layers are needed in order to mitigate the spreading of delamination and obtain realistic delaminated zones. As the laminates are comprised of carbon fiber and epoxy sheets it is of importance to include damage mechanisms inside each individual sheet. Accordingly, a composite material model built into the software is used which considers tensile and compressive stress in fiber and epoxy. The strength limits are then set according to experimental test data. The bird is modeled using a mesh free technique called Smooth Particle Hydrodynamics using a material model with properties similar to a fluid. The internal pressure of the bird model is linked to the change in volume with an Equation of State. By examining the bird models behavior compared to experimental results it is determined to have a realistic impact on structures. A model of the leading edge is then subjected to bird strike according to European standards. The wing skin is penetrated indicating that reinforcements might be needed in order to protect valuable components inside the wing structure such as the fuel tank. However, the results are not completely accurate due to the fact that there is little experimental data available regarding soft body penetration of composite laminates. As a consequence, the simulation cannot be confirmed against real experimental results and further investigations are required in order to have confidence in modeling such events. Furthermore, the delamination due to the bird strike essentially spreads across the whole model. Since only one layer of delamination is included the spread is most likely overestimated.
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11

Dunn, Leigh. "Investigating accidents involving aircraft manufactured from polymer composite materials." Thesis, Cranfield University, 2013. http://dspace.lib.cranfield.ac.uk/handle/1826/8448.

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This thesis looks into the examination of polymer composite wreckage from the perspective of the aircraft accident investigator. It develops an understanding of the process of wreckage examination as well as identifying the potential for visual and macroscopic interpretation of polymer composite aircraft wreckage. The in-field examination of aircraft wreckage, and subsequent interpretations of material failures, can be a significant part of an aircraft accident investigation. As the use of composite materials in aircraft construction increases, the understanding of how macroscopic failure characteristics of composite materials may aid the field investigator is becoming of increasing importance. The first phase of this research project was to explore how investigation practitioners conduct wreckage examinations. Four accident investigation case studies were examined. The analysis of the case studies provided a framework of the wreckage examination process. Subsequently, a literature survey was conducted to establish the current level of knowledge on the visual and macroscopic interpretation of polymer composite failures. Relevant literature was identified and a compendium of visual and macroscopic characteristics was created. Two full-scale polymer composite wing structures were loaded statically, in an upward bending direction, until each wing structure fractured and separated. The wing structures were subsequently examined for the existence of failure characteristics. The examination revealed that whilst characteristics were present, the fragmentation of the structure destroyed valuable evidence. A hypothetical accident scenario utilising the fractured wing structures was developed, which UK government accident investigators subsequently investigated. This provided refinement to the investigative framework and suggested further guidance on the interpretation of polymer composite failures by accident investigators.
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Alonso, Rodolfo Delatore. "Comparison between aluminum and all composite AL5 aircraft fuselage." Instituto Tecnológico de Aeronáutica, 2008. http://www.bd.bibl.ita.br/tde_busca/arquivo.php?codArquivo=1160.

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This work demonstrate the decision that lead the responsible team for the AL5 aircraft to recommend more dedicated study, on using composite materials on the constant section barrel of the fuselage, to the next design team. The recommendation where made after the mass comparison of a composite fuselage and an equivalent aluminum fuselage, both estimated through methodologies described on the books of Bruhn and Niu.
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Mativo, John M. "System Design of Composite Thermoelectrics for Aircraft Energy Harvesting." University of Dayton / OhioLINK, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=dayton1607959975788155.

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14

Watkins, R. I. "Multilevel optimum design of large laminated composite structures." Thesis, Cranfield University, 1986. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.374011.

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15

Mahdi, Stephane. "The performance of bonded repairs to composite structures." Thesis, Imperial College London, 2001. http://hdl.handle.net/10044/1/7815.

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16

Wen, Edward A. "Compressive strength prediction for composite unmanned aerial vehicles." Morgantown, W. Va. : [West Virginia University Libraries], 1999. http://etd.wvu.edu/templates/showETD.cfm?recnum=959.

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Thesis (M.S.)--West Virginia University, 1999.<br>Title from document title page. Document formatted into pages; contains ix, 117 p. : ill. (some col.) Includes abstract. Includes bibliographical references (p. 83-84).
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17

Becerra, Pozo Natalia I. "Analysis and optimisation of composite truss structures for aircraft applications." Thesis, Oxford Brookes University, 2006. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.444336.

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18

Brown, Eric L. (Eric Lee). "Integrated strain actuation in aircraft with highly flexible composite wings." Thesis, Massachusetts Institute of Technology, 2003. http://hdl.handle.net/1721.1/8001.

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Thesis (Sc. D.)--Massachusetts Institute of Technology, Dept. of Mechanical Engineering, 2003.<br>Includes bibliographical references (p. 167-171).<br>The purpose of this thesis is to provide a framework for the study of wing warping as a means of achieving multiple aeroelastic goals. Shape change is achieved by integrating anisotropic piezoelectric composites (APC) within the passive composite wing skin. The goals include the ability of integrated strain actuation (ISA) to provide sufficient wing deformation for roll maneuver, gust load alleviation, flutter suppression, and redistribution of maneuver loads. A nonlinear analysis tool was constructed to study the behavior of aircraft with highly flexible active wings undergoing large deformation. It combines a new large displacement, strain-based finite element beam formulation with a finite-state unsteady aerodynamics model and a thin-wall active beam cross section model. The tool was created with the flexibility to model different aircraft configurations, including unconventional ones such as joined wing designs. The effects of sweep and dihedral, as well as large deformations are taken into account in the calculation of aerodynamic loads. The strain-based finite element formulation allows for a simplified control design because the flexible degrees of freedom are easily accessible by strain gages. To support the evaluation of ISA performance, and to study the impact of vehicle size on performance, three representative conventional vehicles using aileron control are modeled. The vehicles are based on fielded unmanned aerial vehicles (UAV), representing low, medium, and high altitude classes. The ISA wings are modeled by replacing some of the passive composite plies in with APC. The active and passive vehicles are compared based on the goals stated above. The impact of the piezoelectric composite material properties on weight and performance is discussed.<br>(cont.) The required property values are determined for making ISA a viable method for primary roll control and wing stability. Numerical results show that roll control without ailerons is possible using present actuator technology. Integrated strain actuation is also shown to significantly alleviate gust loading and increase the flutter speed. Peak maneuver stresses are significantly reduced through active lift redistribution.<br>by Eric L. Brown.<br>Sc.D.
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Kapidzic, Zlatan. "Strength analysis and modeling of hybrid composite-aluminum aircraft structures." Licentiate thesis, Linköpings universitet, Hållfasthetslära, 2013. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-91894.

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The current trend in aircraft design is to increase the proportion of fiber composites in the structures. Since many primary parts also are constructed using metals, the number of hybrid metal-composite structures is increasing. Such structures have traditionally often been avoided as an option because of the lack of methodology to handle the mismatch between the material properties. Composite and metal properties differ with respect to: thermal expansion, failure mechanisms, plasticity, sensitivity to load type, fatigue accumulation and scatter, impact resistance and residual strength, anisotropy, environmental sensitivity, density etc. Based on these differences, the materials are subject to different design and certification requirements. The issues that arise in certification of hybrid structures are: thermally induced loads, multiplicity of failure modes, damage tolerance, buckling and permanent deformations, material property scatter, significant load states etc. From the design point of view, it is a challenge to construct a weight optimal hybrid structure with the right material in the right place. With a growing number of hybrid structures, these problems need to be addressed. The purpose of the current research is to assess the strength, durability and thermo-mechanical behavior of a hybrid composite-aluminum wing structure by testing and analysis. The work performed in this thesis focuses on the analysis part of the research and is divided into two parts. In the first part, the theoretical framework and the background are outlined.Significant material properties, aircraft certification aspects and the modeling framework are discussed.In the second part, two papers are appended. In the first paper, the interaction of composite and aluminum, and their requirements profiles,is examined in conceptual studies of the wing structure. The influence of the hybrid structure constitution and requirement profiles on the mass, strength, fatigue durability, stability and thermo-mechanical behavior is considered. Based on the conceptual studies, a hybrid concept to be used in the subsequent structural testing is chosen. The second paper focuses on the virtual testing of the wing structure. In particular, the local behavior of hybrid fastener joints is modeled in detail usingthe finite element method, and the result is then incorporated into a global model using line elements. Damage accumulation and failure behavior of the composite material are given special attention. Computations of progressive fastener failure in the experimental setup are performed. The analysis results indicate the critical features of the hybrid wing structure from static, fatigue, damage tolerance and thermo-mechanical points of view.
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Clark, Randal John. "Damage tolerance of bonded composite aircraft repairs for metallic structures." Thesis, University of British Columbia, 2007. http://hdl.handle.net/2429/31275.

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This thesis describes the development and validation of methods for damage tolerance substantiation of bonded composite repairs applied to cracked plates. This technology is used to repair metal aircraft structures, offering improvements in fatigue life, cost, manufacturability, and inspectability when compared to riveted repairs. The work focuses on the effects of plate thickness and bending on repair life, and covers fundamental aspects of fracture and fatigue of cracked plates and bonded joints. This project falls under the UBC Bonded Composite Repair Program, which has the goal of certification and widespread use of bonded repairs in civilian air transportation. This thesis analyses the plate thickness and transverse stress effects on fracture of repaired plates and the related problem of induced geometrically nonlinear bending in unbalanced (single-sided) repairs. The author begins by developing a classification scheme for assigning repair damage tolerance substantiation requirements based upon stress-based adhesive fracture/fatigue criteria and the residual strength of the original structure. The governing equations for bending of cracked plates are then reformulated and line-spring models are developed for linear and nonlinear coupled bending and extension of reinforced cracks. The line-spring models were used to correct the Wang and Rose energy method for the determination of the long-crack limit stress intensity, and to develop a new interpolation model for repaired cracks of arbitrary length. The analysis was validated using finite element models and data from mechanical tests performed on hybrid bonded joints and repair specimens that are representative of an in-service repair. This work will allow designers to evaluate the damage tolerance of the repaired plate, the adhesive, and the composite patch, which is an airworthiness requirement under FAR (Federal Aviation Regulations) 25.571. The thesis concludes by assessing the remaining barriers to certification of bonded repairs, discussing the results of the analysis, and making suggestions for future work. The developed techniques should also prove to be useful for the analysis of fibre-reinforced metal laminates and other layered structures. Some concepts are general and should be useful in the analysis of any plate with large in-plane stress gradients that lead to significant transverse stresses.<br>Applied Science, Faculty of<br>Mechanical Engineering, Department of<br>Graduate
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DeVries, Dean. "Design of a cure monitoring system for composite aircraft repair patches." Thesis, National Library of Canada = Bibliothèque nationale du Canada, 1996. http://www.collectionscanada.ca/obj/s4/f2/dsk3/ftp04/MQ45454.pdf.

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22

Dayyani, Iman. "Mechanical behavior of composite corrugated structures for skin of morphing aircraft." Thesis, Swansea University, 2015. https://cronfa.swan.ac.uk/Record/cronfa42865.

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Corrugated panels have gained considerable popularity in a range of engineering applications, particularly in morphing skin applications due to their remarkable anisotropic characteristics. They are stiff to withstand the aerodynamic loads and flexible to enable the morphing deformations. In this thesis a detailed review of the literature on corrugated structures is presented. The specific characteristics of corrugated structures such as: high anisotropic behaviour, high stiffness and good durability, lightness and cost effectiveness are discussed comprehensively. However for the application in morphing aircraft, the optimal design of the corrugated panels requires simple models of these structures to be incorporated into multi-disciplinary system models. Therefore equivalent structural models are required that retain the dependence on the geometric parameters and material properties of the corrugated panels. In this regard, two analytical solutions based on homogenization and super element techniques are presented to calculate the equivalent mechanical properties of the corrugated skin. Different experimental and numerical models are investigated to verify the accuracy and efficiency of the presented equivalent models. The parametric studies of different corrugation shapes demonstrate the suitability of the proposed super element for application in further detailed design investigations. Then the design and multi-objective optimization of an elastomer coated composite corrugated skin for the camber morphing aerofoil is presented. The geometric parameters of the corrugated skin are optimized to minimize the in-plane stiffness and the weight of the skin and to maximize the flexural out-of-plane stiffness of the corrugated skin. A finite element code for thin beam elements is used with the aggregate Newton's method to optimize the geometric parameters of the coated corrugated panel. The advantages of the corrugated skin over the elastomer skin for the camber morphing structure are discussed. Moreover, a finite element simulation of the camber morphing internal structure with the corrugated skin is performed under typical aerodynamic and structural loadings to check the design approach.
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Aljets, Dirk. "Acoustic emission source location in composite aircraft structures using modal analysis." Thesis, University of South Wales, 2011. https://pure.southwales.ac.uk/en/studentthesis/acoustic-emission-source-location-in-composite-aircraft-structures-using-modal-analysis(6871e94b-6e94-4efd-b563-41b254ee27b4).html.

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The aim of this research work was to develop an Acoustic Emission (AE) source location method suitable for Structural Health Monitoring (SHM) of composite aircraft structures. Therefore useful key signal features and sensor configurations were identified and the proposed method was validated using both artificially generated AE as well as actual AE resulting from damage. Acoustic Emission is a phenomenon where waves are generated in stressed materials. These waves travel through the material and can be detected with suitable sensors on the surface of the structure. These stress waves are attributed to propagating damage inside the material and can be monitored while the structure is in service. This makes AE very suitable for SHM, in particular for aircraft structures. In recent years composite materials such as carbon fibre reinforced epoxy (CFRP) are increasingly being used for primary and secondary structures in aircraft. The anisotropic layup of CFRP can lead to different failure mechanisms such as delamination, matrix cracking or fibre breakage which affects the remaining life time of the structure to different extents. Accurate damage location is important for SHM systems to avoid further inspections and allows for a maintenance scheme which considers the severity of the damage, due to damage type, extent and location. This thesis presents a novel source location method which uses a small triangular AE sensor array. The method determines the origin of an AE wave by a combination of time of arrival and modal analysis. The small footprint of the array allows for a fast and easy installation in hard-to-reach areas. The possibility to locate damage outside and at a relatively far distance from the array could potentially reduce the overall number of sensors needed to monitor a structure. Important wave characteristics and wave propagation in particular in CFRP were investigated using AE simulated by an artificial source and actual damage in composite specimens.
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Davies, Thomas Gethin. "The analysis of bonded repair solutions for primary composite aircraft structures." Thesis, Swansea University, 2013. https://cronfa.swan.ac.uk/Record/cronfa42862.

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Svalstedt, Mats, and Sofia Swedberg. "Commercial Aircraft Wing Structure : - Design of a Carbon Fiber Composite Structure." Thesis, KTH, Skolan för teknikvetenskap (SCI), 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-276702.

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This project explores the classical wing structure of an commercial aircraft for an all carbon fiber reinforced polymer unmanned aerial vehicle(UAV). It is part of a collaborative work consisting of several groups researching different parts of the aircraft. The objective of this report is to present the design of the inner wing structure for a greener, more efficient scaled 2:1 version of the Skywalker X8. In order to make the aircraft as efficient as possible, the structure needs to be lightweight. The loads were first approximated using XFLR5 and a first design made. The design was then tested using finite element analysis (FEA) in the programme Ansys Static Structural. The material that was tested was carbon fiber/epoxy prepreg. The final design of the wing weighs 3.815 kg, and consists of one spar and a skin thickness of 1 mm. The weight of the whole aircraft, including the propulsion system and a sharklet at both wingtips researched by other groups, is 20.262 kg. The lift-to-drag ratio was also calculated, and the most efficient angle of attack was concluded to be around 2-3°.<br>Detta projekt utforskar den klassiska vingstrukturen av ett kommersiellt flygplan för en obemannad luftfarkost gjord helt i kolfiberarmerad polymer. Det är en del av ett samarbete som består av flera projektgrupper som forskar på olika delar av flygplanet. Målet med projektet är att designa den inre vingstrukturen för en miljövänligare, mer effektiv uppskalad 2:1 version av drönaren Skywalker X8. För att göra flygplanet så effektiv som möjligt så behöver den vara lättviktig. Lasterna var först uppskattade via XFLR5 och en första design gjordes. Designen testades sedan med finita elementmetoden (FEM) i programmet Ansys Static Structural. Materialet som testades var kolfiber/epoxi prepreg. Den slutgiltiga vingdesignen väger 3.815 kg, och består av en bom och en tjocklek på 1 mm av vingskalet. Totala vikten av flygplanet, inklusive framdrivningssystemet samt virveldämpare på båda vingspetsarna som är framtagna av andra grupper, är 20.262 kg. Glidtalet beräknades även, och är som mest effektiv runt 2-3°.
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Armstrong, Keith Bernard. "The selection of adhesives and composite matrix resins for aircraft repairs." Thesis, City, University of London, 1990. http://openaccess.city.ac.uk/18929/.

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This Thesis studies various aspects of the selection of adhesives for aircraft repairs. It was undertaken because of the need to repair a number of different aircraft types with a limited range of adhesives to minimise stockholdings and wastage of materials which have shelf-life limitations. The aspects studied were tensile strength, tensile modulus, elongation at failure, fracture energy, compression strength, compression modulus, water uptake, Tg dry and we t and diffusion coefficient. A computer programme written at Leicester Polytechnic was adapted to suit an IBM PC and this allowed water uptake within a lap joint to be related to time, diffusion coefficient and solubility coefficient. Later in the project attempts were made to relate the tensile properties to lap joint strength. They were partially successful and led to a number of wedge tests to obtain fracture energy data. From this fracture toughness was calculated and this gave the best correlation with lap joint strength. It was finally concluded that good lap joint strength required an optimum combination of tensile strength, modulus, elongation to failure and fracture energy to achieve a fracture toughness of at least 3 MN.m-3/2 The fracture energy data clearly separated the brittle composite "matrix resins" from the tougher "adhesives" • It was concluded that matrix resins and adhesives can be more easily and effectively compared using the fundamental properties of the resins themselves than by using the limited data normally supplied by the Manufacturers on their data sheets. It was also considered that the simpler tensile, compression and wedge tests used in this programme could obtain more data, more quickly and more cheaply than the thick adherend or napkin ring tests usually advocated. From the results obtained it was possible to suggest specifications for both composite matrix resins and adhesives for making bonded metal or composite joints. These should lead to the development of better matrix resins and adhesives.
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Soykasap, Omer. "Aeroelastic optimization of a composite tilt rotor." Diss., Georgia Institute of Technology, 1999. http://hdl.handle.net/1853/11823.

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Carmai, Julaluk. "The modelling of matrix-coated fibre composite consolidation." Thesis, University of Oxford, 2001. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.365724.

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Ashok, Kumar Sachin Sharma. "Incorporation of graphene thin films into the carbon fiber reinforced composite via 3d composite concept against the lightning strikes on composite aircraft." Thesis, Wichita State University, 2012. http://hdl.handle.net/10057/5592.

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Research and development of graphene and graphene based materials have been increasing significantly since they were invented. This report presents the development of a highly conductive graphene thin film (GTF) to reduce the damage of lightning effects on composite aircrafts. Furthermore, there are three new developments that are presented in this research: (a) the development of a highly conductive functionalized nanosize GTF, (b) a new approach of incorporating the GTF into the carbon fiber reinforced composite panel, and (c) a new development of 3D stitching concept were introduced specifically using polyester threads instead of fiber yarns that can be useful for the applications of aircraft protections against the effects of lightning strike. In addition, graphene was chemically functionalized and oxidized to form GTF. The highest electrical conductivity measured on the GTF was approximately 1800 S/cm. Furthermore, the GTF was then incorporated into the carbon fiber reinforced composite. Delamination was observed between the GTF and the composite. To investigate this issue, the composite was mechanically tested and there was a 40% decrease in tensile strength compared to the baseline. Therefore, 3D stitching concept was then introduced to reduce the delamination. Four stitch configurations having different stitch length, thread to thread thickness, thread tension, and thread thickness respectively were used in this study. 3D stitching was initially done on six sheets of unidirectional prepreg, MTM45-1 without the incorporation of GTF. Furthermore, mechanical testing was carried out and the stitch configuration that delivered the most appropriate result was then further used on twelve sheets of unidirectional prepreg, 5320. Here, three samples which was the baseline, 5320 with and without the incorporation of GTF respectively were prepared and mechanically tested, and the strength values were observed.<br>Thesis (M.S.)--Wichita State University, College of Engineering, Dept. of Mechanical Engineering
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Liu, Hongfen. "A structural design comparison of metallic and composite aircraft pressure retaining doors." Thesis, Cranfield University, 2012. http://dspace.lib.cranfield.ac.uk/handle/1826/7308.

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The pressure retaining door is obviously a sensible part of an aircraft, and the design criteria is much more critical than for the fuselage, so a problem caused by this critical criteria is the heavy weight of the door structure because it should be strong enough to withstand loads and stiff enough to meet the sealing requirements. In spite of the pressure retaining door being so important, it is difficult to find design references. So, in this thesis, the pressure retaining door is investigated first, and then a typical structure of a type A door is selected as the study case using both metallic and composite material, in order to generate a standard method for door structure design, and to identify the key factors which can affect the structure weight. The study indicates that the structure weight of a type A door can be kept in a range for different combinations of beams and stringers, and the composite door structure can be 20% lighter than the metallic door while the stiffness of the two doors remains similar. It is found that the skin contributes much more weight to the door structure than other components and the skin thickness is affected by the short edge of the skin panel divided by beams and stringers. The results also found that it is much more serious when the end stop fails than when the middle stops fail. Therefore, it appears that the composite door is a good material as an alternative to aluminium. Also the method of door structure design is reasonable for the composite door, although it would be better to consider the stiffness of beams while in the theory design period. Besides IRP, the Group Design Project (GDP) is another important part of the MSc study; it lasts nearly half a year and we complete the Fly-wing concept design. The main contribution of the author to the GDP is the arrangement of doors, and also includes the family issues, cabin layout arrangement and a 3D model construct, which can be seen in APPENDIX B. According to the GDP work, I will have broadened my professional knowledge and will have an overall view of aircraft design.
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Bail, Justin L. "Non-desctructive investigation & FEA correlation on an aircraft sandwich composite structure." Akron, OH : University of Akron, 2007. http://rave.ohiolink.edu/etdc/view?acc%5Fnum=akron1196702586.

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Thesis (M.S.)--University of Akron, Dept. of Civil Engineering, 2007.<br>"December, 2007." Title from electronic thesis title page (viewed 02/25/2008) Advisor, Wieslaw Binienda; Faculty readers, Craig Menzemer, Robert Goldbert; Department Chair, Wieslaw Binienda; Dean of the College, George K. Haritos; Dean of the Graduate School, George R. Newkome. Includes bibliographical references.
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Eslimy-Isfahany, Seyed Hamid Reza. "Dynamic response of thin-walled composite structures with application to aircraft wings." Thesis, City University London, 1998. http://openaccess.city.ac.uk/7719/.

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A general analytical method is developed to study first the buckling behaviour and then the dynamic characteristics of thin-walled composite structures with the presence of bending torsion coupling. The dynamic response theory incorporates the dynamic stiffness matrix approach and generalised coordinates using the normal mode method. Structural components considered are thin-walled laminated composite beams with carbon-fibre, glass-fibre or other reinforced plastic lay-ups. The examples of such beams and their applications include aircraft wings, hulls of ships, helicopter and wind turbine blades. All assumptions made in this work are based on elastic linear small deflection beam theory so that the overall response of the beam is represented by the superposition of all individual responses in each mode. Bending-torsion coupling effects arising from the anisotropic nature of fibrous composites, as well as due to non-coincident centroid and geometric shear centre of the beam crosssection, are the main contributory elements when developing the theory. The beam is subjected to time dependent forces and/or torques which can be either concentrated or distributed over its length. Both deterministic and random loads are considered. An important example of a deterministic load is one that varies harmonically in time. The Duhamel integral is employed to calculate the response to any arbitrary time dependent deterministic load. The random load is assumed to be Gaussian, having both stationary and ergodic properties. The evaluation of the response to the random load is carried out in the frequency domain by relating the Power Spectral Density (PSD) of the output to that of the input using the complex frequency response function. A number of PSD distributions are considered as random input in order to determine the PSD of the dynamic response. Atmospheric turbulence, which is considered to be one of the forms of random excitation, is modelled using the von Karman spectra for composite aircraft wings. In order to establish the methodology, bending-torsion coupled metallic beams are first ,investigated. The bending-torsion coupling in such beams occurs due to non-coincident centroid and geometric shear centre of the beam cross-section. The natural frequencies and mode shapes in undamped free vibration are obtained and the significance of generalised ,mass in each of the modes of vibration is evaluated. A normal mode method is then used to compute the frequency response function of the beam. The effects of shear deformation rotatory inertia and axial load on the frequencies, mode shapes and dynamic response characteristics are demonstrated. It was essential at an earlier stage of the investigation to validate the chosen composite beam modelling. Among all the different techniques used to determine the rigidities of a composite beam, the buckling load provides a reasonable estimate. The elastic critical buckling loads of thin-walled laminated composite columns for various end conditions are established theoretically using the exact stiffness method. The effect of shear deformation on the buckling characteristics of the column is demonstrated. Experiments are carried out to establish the elastic critical buckling load of metallic and laminated composite columns. Theoretical predictions of the buckling behaviour are corroborated by experimental results and other published results. The investigation is then focused on composite beams, but the response analysis of such beams is significantly more complicated than that of their metallic counterparts. This is mainly due to anisotropic characteristics of laminated fibrous composites. A detailed parametric study with the variation of significant composite parameters, such as ply angle, is undertaken and the importance of the results are highlighted. A suite of computer programs in FORTRAN is developed to predict the bucklingbehaviour, the free vibration and the responsec characteristics of thin-walled composite or metallic beams based on the theory proposed. Numerical results are presented, fully discussed and commented on.
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Benchekchou, Boutaina. "Stresses around fasteners in composite aircraft structures and effects on fatigue life." Thesis, University of Southampton, 1994. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.241160.

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Land, Ian B. (Ian Brett). "Design and manufacture of advanced composite aircraft structures using automated tow placement." Thesis, Massachusetts Institute of Technology, 1996. http://hdl.handle.net/1721.1/31076.

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Thesis (S.M.)--Massachusetts Institute of Technology, Sloan School of Management; and, Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Mechanical Engineering, 1996.<br>Includes bibliographical references (leaves 89-91).<br>by Ian B. Land.<br>S.M.
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Bail, Justin. "Non-Destructive Investigation & FEA Correlation on an Aircraft Sandwich Composite STructure." University of Akron / OhioLINK, 2007. http://rave.ohiolink.edu/etdc/view?acc_num=akron1196702586.

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36

Ellwood, Jeffrey L. "Design and construction of a composite airframe for UAV research." Thesis, Monterey, California : Naval Postgraduate School, 1990. http://handle.dtic.mil/100.2/ADA232422.

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Thesis (M.S. in Aeronautical Engineering)--Naval Postgraduate School, June 1990.<br>Thesis Advisor(s): Howard, Richard M. Second Reader: Lindsey, Gerald H. "June 1990." Description based on signature page as viewed on October 21, 2009. DTIC Identifier(s): Composite materials, ducted fan, airframes, vertical takeoff aircraft, remotely piloted vehicles. Author(s) subject terms: UAV, composites, AROD, TDF, RPV, ducted fan, vertical takeoff. Includes bibliographical references (p. 74-75). Also available online.
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Evans, Simon. "Characterisation of outgassing from carbon fibre composite aircraft joints subjected to lightning current." Thesis, Cardiff University, 2018. http://orca.cf.ac.uk/117887/.

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Outgassing is a lightning direct effect that may occur at structural joints in the fuel laden volume of an aircraft. If uncontrolled, the event is extremely hazardous due to its potential to cause fuel vapour ignition. The aerospace industry has been aware of the threat for many years and lightning strike protection is well established. However, there is a lack of understanding particularly concerning the fundamental mechanisms for the creation of the event. Modern aircraft designs that utilise materials such as carbon fibre reinforced plastic (CFRP), are more dependent on manufacturing process control. Knowledge of the fundamental mechanisms responsible for outgassing can enable relaxation of specifications concerning manufacturing variables that exist, specifically, for the lightning protection of CFRP structures. Evidence from previous studies has revealed the significance of parameters relating to the interface between the fastener and the surrounding structure. However, the electrical parameter that drives the creation of the phenomenon remains unclear. The principle aim of this thesis was to determine a single measurable electrical parameter related to outgassing intensity in CFRP structures which can be used as a performance metric for the optimisation of lightning strike protection. Following the execution of three controlled experiments, it was found that outgassing intensity is a direct consequence of the magnitude of electrical energy absorbed, specifically, at the interface between the fastener shank and the surrounding CFRP structure. Characterisation techniques for robust voltage measurement and the distribution of current to the critical interface were developed to deduce the magnitude of energy absorption. This critical parameter can be now used as a control parameter in future aircraft development for the optimisation of lightning strike protection.
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Kapidzic, Zlatan. "Static and Fatigue Failure of Bolted Joints in Hybrid Composite-Aluminium Aircraft Structures." Doctoral thesis, Linköpings universitet, Mekanik och hållfasthetslära, 2015. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-122349.

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The use of fibre composites in the design of load carrying aircraft structures has been increasing over the last few decades. At the same time, aluminium alloys are still present in many structural parts, which has led to an increase of the number of hybrid composite-aluminium structures. Often, these materials are joined at their interface by bolted connections. Due to their different response to thermal, mechanical and environmental impact, the composite and the aluminium alloy parts are subject to different design and certification practices and are therefore considered separately.The current methodologies used in the aircraft industry lack well-developed methods to account for the effects of the mismatch of material properties at the interface.One such effect is the thermally induced load which arises at elevated temperature due to the different thermal expansion properties of the constituent materials. With a growing number of hybrid structures, these matters need to be addressed.  The rapid growth of computational power and development of simulation tools in recent years have made it possible to evaluate the material and structural response of hybrid structures without having to entirely rely on complex and expensive testing procedures.However, as the failure process of composite materials is not entirely understood, further research efforts are needed in order to develop reliable material models for the existing simulation tools. The work presented in this dissertation involves modelling and testing of bolted joints in hybrid composite-aluminium structures.The main focus is directed towards understanding the failure behaviour of the composite material under static and fatigue loading, and how to include this behaviour in large scale models of a typical bolted airframe structure in an efficient way. In addition to that, the influence of thermally induced loads on the strength and fatigue life is evaluated in order to establish a design strategy that can be used in the industrial context. The dissertation is divided into two parts. In the first one, the background and the theory are presented while the second one consists of five scientific papers.
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Crump, Duncan Andrew. "Performance analysis of a reduced cost manufacturing process for composite aircraft secondary structure." Thesis, University of Southampton, 2009. https://eprints.soton.ac.uk/142803/.

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In the current, environmentally-aware, climate aircraft designers are under increasing pressure to produce fuel efficient vehicles. Weight reduction is an important method for increasing fuel efficiency. Fibre reinforced polymer (FRP) composites are known to offer weight savings over traditional metallic components, due to their excellent stiffness and strength to weight ratios. However, the major limiting factor for the use of aerospace quality composites is the manufacturing cost. The costs incurred in the conventional process of prepreg cured in an autoclave are well documented. The research in this thesis is concerned with reducing the cost of manufacturing aircraft standard carbon fibre composite sandwich panels, whilst maintaining mechanical performance. The overall aim of the EngD is to provide a unified approach for assessing the performance of carbon fibre sandwich secondary structure that are manufactured using several different techniques. Cost and performance criteria are defined so that an optimal panel can be produced. The work has been motivated by the industrial sponsor, GE Aviation Systems. Five combinations of raw material and processing techniques, manufacturing options (MOs) were considered in incremental steps from the baseline of unidirectional prepreg cured in an autoclave to the noncrimp fabric (NCF) infiltrated using resin film infusion (RFI) and cured in a conventional oven. For cost and performance analysis a generic panel has been designed that is representative of secondary wing structure on commercial passenger aircraft. The cost was estimated by monitoring the manufacture of generic panels using each MO, whilst the performance was measured by both mechanical characterisation tests and by full scale tests on a custom designed rig. The rig applies a pressure load using a water cushion and allows optical access to the surface of the panel enabling the use of optical techniques, i.e. thermoelastic stress analysis (TSA) and digital image correlation (DIC). Feasibility tests on TSA and DIC demonstrated their use on the materials considered in this thesis, and were used to validate finite element (FE) models. The RFI out-of-autoclave process was found to reduce generic panel manufacture time by almost 30%, and the material cost was reduced by almost 40%. The mechanical characterisation tests suggested the ‘new’ process could produce laminates with a similar fibre volume fraction to that of the original process and similar in and out-of-plane mechanical properties. The in-plane stiffness was slightly reduced by 7 %, but the strength showed an increase of 12%. Full scale tests on the generic panels using point out-of-plane deflection measurements and full field TSA demonstrated the panel produced using the ‘new’ process has adequate performance. Moreover the full-field tests indicated an improvement in performance. Further work is required to optimise the design of the panel for weight, in particular the weight of the raw material, and investigating methods for modelling the NCF for certification.
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40

Morris, Russell A. "Aeroelastic modeling and flutter control in aircraft with low aspect ratio composite wings." Diss., Virginia Tech, 1996. http://hdl.handle.net/10919/40462.

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41

Foti, Federico. "Effect of the Environment on the Fatigue Behaviour of Textile Organic Matrix Composite Materials for Aircraft Applications." Thesis, Chasseneuil-du-Poitou, Ecole nationale supérieure de mécanique et d'aérotechnique, 2017. http://www.theses.fr/2017ESMA0031/document.

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Les composites à matrice organiques (CMO) et fibre de carbone sont de plus en plus employés dans la réalisation de structures « tièdes » (aubes de fan, nacelles …) ; ces pièces peuvent être soumises, en service, à la fatigue mécanique, au cyclage thermique et à la fatigue thermo-mécanique. Bien qu’il existe une littérature consistante sur le comportement en fatigue des composites tissés, l'interaction entre fatigue et la dégradation liée à l'environnement à haute température n’a pas été encore bien exploitée. Le couplage entre les effets de la thermo-oxydation, le comportement mécanique (viscoélastique, viscoplastique) de la matrice organique à températures élevées et la dégradation par fatigue peut être néfaste pour le composite.Le but de ce travail est de caractériser et de modéliser - pour les composites tissés C/matrice organique - le comportement thermomécanique, l'apparition et le développement de l’endommagement liés aux mécanismes mécaniques cycliques (fatigue) sous environnement contrôlé (température et gaz).Une étude préliminaire sur un composite stratifiée [02/902]s a été menée pour pouvoir analyser les effets de l’environnement sur une architecture simple. La corrélation d’image numérique (CIN) et des scans μ-tomographiques (μCT) ont été employés pour le suivi et la caractérisation de l’endommagement de fatigue de composites tissés 2D à architecture complexe pour applications aéronautiques. Les effets de l’environnement sur la dégradation par fatigue ont été également explorés.L'objectif à long terme de cette étude est de fournir des outils expérimentaux et numériques pour renforcer la compréhension et la modélisation du couplage mécanique/endommagement/environnement pour la prédiction de la durée de vie et pour la proposition de protocoles d’essais accélérés réalistes de pièces « tièdes » en CMO<br>In the next future, the employment of organic matrix/carbon fibre composites (OMC) is foreseen for the realization of “hot” structures: these parts may be subjected, in service, to mechanical fatigue (e.g. fan blades turbo-engines), thermal cycling and thermo-mechanical fatigue (e.g. aircraft structural parts). Though there is a consistent literature concerning the fatigue behaviour of woven composites, the interaction between fatigue and environmental degradation at high temperature has been poorly explored. Coupling between thermo-oxidation effects, mechanical (viscoelastic, viscoplastic) behaviour of the polymer matrix at high temperatures and degradation due to fatigue may be highly detrimental for the material. This work aims at characterizing and modelling - for carbon fibre/organic matrix (polyimide) textile composites – the thermomechanical behaviour, the onset and the development of damage related to cyclic mechanical mechanisms (fatigue) under controlled (temperature and gas) environment.A preliminary study on a cross-ply laminate [02/902]s has been carried out in order to analyse the environmental effect on a model sample. Digital Image Correlation (DIC) and μ-Computed Tomography (μCT) have been used to monitor and characterize the fatigue damage of 2D woven composites for aeronautical applications. The environmental effect on fatigue degradation have been also explored.The long-term aim of the study is to provide experimental and numerical tools to strengthen the understanding and the modelling of mechanics/damage/environment coupling for durability prediction
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42

Kaufmann, Markus. "Cost/Weight Optimization of Aircraft Structures." Licentiate thesis, Stockholm : Farkost- och flyg, Kungliga Tekniska högskolan, 2008. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-4645.

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43

Farrow, I. R. "Damage accumulation and degradation of composite laminates under aircraft service loading : assessment and prediction." Thesis, Cranfield University, 1989. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.237804.

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44

Rudd, Jeffrey Roy. "COMPRESSIVE STRENGTH TO WEIGHT RATIO OPTIMIZATION OF COMPOSITE HONEYCOMB THROUGH ADDITION OF INTERNAL REINFORCEMENTS." University of Akron / OhioLINK, 2006. http://rave.ohiolink.edu/etdc/view?acc_num=akron1145900147.

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45

Rubenking, Samuel Kim. "Dual Mode Macro Fiber Composite-Actuated Morphing Tip Feathers for Controlling Small Unmanned Aircraft." Thesis, Virginia Tech, 2017. http://hdl.handle.net/10919/78433.

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The transition of flight from manned to unmanned systems has led to new research and applications of technology within the field that, until recently, were previously thought to be unfeasible. The industry has become interested in alternative control surfaces and uses for smart materials. A Macro Fiber Composite (MFC), a smart material, takes advantage of the piezoelectric effect and provides an attractive alternative actuator to servos in the Small Unmanned Aerial Systems (SUAS) regime of flight. This research looks to take MFC actuated control surfaces one step further by pulling inspiration from and avian flight. A dual mode control surface, created by applying two sets of two MFCs to patch of carbon fiber, can mimic the tip feathers of a bird. This actuator was modeled both using Finite Element Analysis (FEA) and Computational Fluid Dynamics (CFD). Real-world static testing on a feather confirmed preliminary FEA results, and wind tunnel tests simulating assumed cruise conditions confirmed the feather would not exhibit any adverse structural behaviors, such as flutter or aeroelastic divergence. From its modeled performance on a wing using CFD, the MFC feather proved to be a success. It was able to produce a wing that, when compared to a traditional rectangular wing, yielded 73% less induced drag and generated proverse yaw. However, the MFC feathers alone, in the configuration tested, did not produce enough roll authority to feasibly control an aircraft.<br>Master of Science
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46

Backhouse, R. "Multiaxial non-crimp fabrics : characterisation of manufacturing capability for composite aircraft primary structure applications." Thesis, Cranfield University, 1998. http://dspace.lib.cranfield.ac.uk/handle/1826/1929.

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Carbon composite reinforcement fabrics aimed at flight critical aircraft structure application were designed and the capability of the process used to manufacture them examined. Studies of the LIBA multiaxial non-crimp fabric manufacturing process focused on the effect of changes to four manufacturing parameters using an experimental design process to design the fabrics and analyse the results. The composite properties measured included microstructural features of the fibre tows and resin distribution, and mechanical performance both in-plane and their damage resistance and tolerance characteristics. Nine pairs of Toray T300 carbon based LIBA multiaxial non-crimp fabrics were manufactured and converted to composite laminates. Processing was accomplished using the interleaved Resin Film Infusion processing route with commercial Fiberdux 914 matrix resin. All the fabrics were of the same reinforcement type, consisting of 816 g/m2 of fibre; 376 g/m2 oriented along the fabric length (0°) and 220 g/m2 oriented in each of the ±45° directions. Differences between the nine pairs of fabrics were restricted to the settings of four manufacturing parameters; stitch course (needle penetrations/cm); stitch tension, 00 tension and 0° coverage (amount of constraint on the 0° material provided by the stitch). Three settings were used for each of the parameters; each representing the upper and lower limits, and standard setting. Microstructural characterisation of the laminates indicated large differences in both resin distribution and levels of 0° fibre crimp caused by the changes in manufacturing parameter settings. In-plane and damage resistance and tolerance tests on their composites allowed relationships between manufacturing settings, microstructure and engineering properties to be deduced. It was found that selected in-plane properties could be increased by as much as 17% relative to standard production materials, although a wide range of influence was observed. For damage resistance and tolerance characteristics, reductions in impact damage area (C-scan) of between 13-50% are expected across a range of energies. Manufacturing settings to maximise the impact force for delamination initiation were found to minimise the impact damage areas. Similarly the same settings maximised both the Mode I propagation strain energy release rate and the Compression After Impact strength of the materials. It was found that polyester knitting yarn was largely responsible for the control of the damage resistance and tolerance characteristics together with the mean size of the resin areas and layers within the composite. The manufacturing/microstructure/property relationships identified provide those wishing to exploit these materials with design guidelines to tailor fabric structure and performance characteristics for the intended application. Above all else the results highlight the need for precision in specifying and controlling the manufacturing process in order to repeatably produce the desired performance. Further work on the same materials could be used to provide a link to processing characteristics such as permeability for liquid resin moulding processes and ability to conform to complex curved surfaces.
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47

Satterwhite, Matthew Ryan. "Development and Validation of Fluid-Structure Interaction in Aircraft Crashworthiness Studies." Thesis, Virginia Tech, 2013. http://hdl.handle.net/10919/51559.

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Current Federal Aviation Regulations require costly and time consuming crashworthiness testing to certify aircraft. These tests are only capable of a limited assessment of progressive damage and all crash configurations and scenarios cannot be physically evaluated. Advancements in technology have led to accurate and effective developments in numerical modeling that have the possibility of replacing these rigorous physical experiments. Through finite element analysis, an in-depth investigation of an aircraft equipped with a fabricated composite undercarriage was evaluated during water ditching. The severe impact of aircraft ditching is dynamic and nonlinear in nature; the goal of this work to develop a methodology that not only captures the structural response of the aircraft, but also the fluidic behavior of the water. Fundamental studies were first conducted on a well-researched fluid-solid interaction problem, the water entry of a wedge. Typical modeling strategies did not capture the desired detail of the event. An advanced meshing scheme combining meshed and meshless Lagrangian techniques was developed and multiple wedge angles were tested and compared to analytic and qualitative results. The meshing technique proved valid, as the difficult to model phenomena of splashing was captured and the maximum impact force was within five percent of analytical calculations for the 20° and 30° deadrise wedge. Physical small scale aircraft ditching experiments were then performed with an innovative testing platform capable of producing varied aircraft approach configurations. The model was outfitted with an instrumented composite undercarriage to record data throughout the impact while a high-speed camera recorded the event. Numerical simulations of the model aircraft were then compared to experimental results with a strong correlation. This methodology was then ultimately tested on a deformable model of a fuselage section of a full-size aircraft.<br>Master of Science
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48

Doepke, Edward Brady. "Design Demonstration and Optimization of a Morphing Aircraft Control Surface Using Flexible Matrix Composite Actuators." Diss., Virginia Tech, 2018. http://hdl.handle.net/10919/82494.

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The morphing of aircraft wings for flight control started as a necessity for the Wright Brothers but quickly fell out of favor as aircraft increased speed. Currently morphing aircraft control is one of many ideas being explored as we seek to improve aircraft efficiency, reduce noise, and other alternative aircraft solutions. The conventional hinged control surface took over as the predominant method for control due to its simplicity and allowing stiffer wings to be built. With modern technologies in variable stiffness materials, actuators, and design methods, a morphing control surface, which considers deforming a significant portion of the wing's surface continuously, can be considered. While many have considered morphing designs on the scale of small and medium size UAVs, few look at it for full-size commercial transport aircraft. One promising technology in this field is the flexible matrix composite (FMC) actuator. This muscle-like actuator can be embedded with the deformable structure and unlike many other actuators continue to actuate with the morphing of the structure. This was demonstrated in the FMC active spoiler prototype, which was a full-scale benchtop prototype, demonstrated to perform under closed-loop control for both the required deflection and load cases. Based on this FMC active spoiler concept a morphing aileron design was examined. To do this an analysis coupling the structure, fluid, and FMC actuator models was created. This allows for optimization of the design with the objectives of minimizing the hydraulic energy required and mass of the system by varying the layout of the FMC aileron, the material properties used, and the actuator's design and placement with the morphing section. Based on a commercial transport aircraft a design case was developed to investigate the optimal design of a morphing aileron using the developed analysis tool. The optimization looked at minimizing the mass and energy requirements of the morphing aileron and was subject to a series of constraints developed from the design case and the physical limitations of the system. A Pareto front was developed for these two objectives and the resulting designs along the Pareto front explored. From this optimization, a series of design guidelines were developed.<br>Ph. D.
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49

Wang, Zhe. "Effects of anti-oxidant migration on friction and wear in carbon-carbon composite aircraft brakes /." Available to subscribers only, 2006. http://proquest.umi.com/pqdweb?did=1240703191&sid=6&Fmt=2&clientId=1509&RQT=309&VName=PQD.

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Thesis (Ph.D.)--Southern Illinois University Carbondale, 2006.<br>"Department of Mechanical Engineering and Energy Processes." Includes bibliographical references (leaves 77-81). Also available online.
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50

Khan, Jehan Zeb. "Static, dynamic and aeroelastic behaviour of thin-walled composite structures with application to aircraft wings." Thesis, City University London, 1992. http://openaccess.city.ac.uk/7992/.

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Theoretical and experimental investigations of the static and dynamic behaviour of thin-walled structures are carried out with the ultimate aim of improving prediction procedures for various aeroelastic phenomena. The dynamic stiffness matrix approach is used for structural idealization, while strip theory and Theodorsen's function C(k) are used for the aerodynamic idealization. The dynamic composite beam with with an axial load centroid, has been carried out using Special cases, that been identified and stiffness matrix for a thin-walled geometric and material coupling together (compressive or tensile) applied at the developed. An exact analysis was then the derived dynamic stiffness matrix. are derivatives of the general case have discussed. A three stage program was developed to compute various static and dynamic properties of thin-walled closed or open section composite beams. In the first stage, equivalent elastic constants (overall laminate moduli) were evaluated for a given stacking sequence and material properties. In the second stage, various sectional properties were computed. When the outputs from these two stages were combined, valuable data on sectional rigidities, mass per unit length, polar mass moment of inertia, and shear centre location from the centroid were obtained. In the third stage of the program, all these properties were used to compute the natural frequencies and normal mode shapes of thin-walled composite structures. These programs can be used individually as well as in a combined manner. An experimental investigation of composite thin plates with varying degrees of bending-torsion coupling was conducted. Flexural and torsional rigidities, natural frequencies, normal mode shapes and flutter speed and frequency were experimentally determined. The results obtained were in close agreement with the theoretical predictions. Various open composite sections were experimentally studied for their static and dynamic properties. The results demanded a more refined investigation of the theory. In addition to the experimental study of composite open sections, a parametric study of uncoupled and coupled frequencies of such sections with common boundary conditions was also conducted. Thin-walled closed aerofoil shaped cantilevered structures were tested to establish flexural and torsional rigidities, shear centre, and the polar-mass-moment of inertia. Natural frequencies and normal mode shapes were also determined. The aeroelastic behaviour of these sections was investigated to establish divergence and flutter characteristics. Comparisons of the experimental results with theoretical predictions of flutter speed and frequency were in general satisfactory and the results provided an insight into the aeroelastic behaviour of thin-walled composite beams. The results are discussed and commented on.
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