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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'endommagementAs 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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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°.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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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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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."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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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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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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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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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.Master of Science
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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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Xu, Rongxin. "Optimal design of a composite wing structure for a flying-wing aircraft subject to multi-constraint." Thesis, Cranfield University, 2012. http://dspace.lib.cranfield.ac.uk/handle/1826/7290.
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This thesis presents a research project and results of design and optimization of a composite wing structure for a large aircraft in flying wing configuration. The design process started from conceptual design and preliminary design, which includes initial sizing and stressing followed by numerical modelling and analysis of the wing structure. The research was then focused on the minimum weight optimization of the /composite wing structure /subject to multiple design /constraints. The modelling, analysis and optimization process has been performed by using the NASTRAN code. The methodology and technique not only make the modelling in high accuracy, but also keep the whole process within one commercial package for practical application.
The example aircraft, called FW-11, is a 250-seat commercial airliner of flying wing configuration designed through our MSc students Group Design Project (GDP) in Cranfield University. Started from conceptual design in the GDP, a high-aspect-ratio and large sweepback angle flying wing configuration has been adopted. During the GDP, the author was responsible for the structural layout design and material selection. Composite material has been chosen as the preferable material for both the inner and outer wing components. Based on the derivation of structural design data in the conceptual phase, the author continued with the preliminary design of the outer wing airframe and then focused on the optimization of the composite wing structure. Cont/d.
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Coleman, Robert Mark 1962. "The effects of design, manufacturing processes and operations management on the assembly of aircraft composite structure." Thesis, Massachusetts Institute of Technology, 1991. http://hdl.handle.net/1721.1/42495.
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Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1991 and Thesis (M.S.)--Sloan School of Management, 1991.Includes bibliographical references (leaf 104).
by Robert Mark Coleman.
M.S.
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Marengo, Giovanni. "The use of unidirectional carbon fibre rods in high loaded joints for a composite large civil aircraft wing structure." Thesis, Cranfield University, 2002. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.393700.
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Pecorella, Daniele. "Methodology for the design and optimization of a morphing wing droop-nose structure for greener aircraft." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2022.
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Droop-Nose Leading Edge (DNLE) morphing wings are one of the most promising devices in order to achieve aerodynamic drag and noise reduction during take-off and landing phases. An accurate design of these structures could lead to the decrease of aircraft fuel consumption in the perspective of reaching a greener aviation, following the objectives indicated by Flightpath 2050 issued by the E.U. However, due to the challenges related to the realization of this technology and TRL reached, DNLE are more likely implemented in Unmanned Aerial Systems (UAS) for testing and evaluation purposes. In the present study, an optimization methodology for the DNLE composite laminate skin and morphing mechanism structure is proposed and applied to a study case represented by the UAS-S45 aircraft. The work starts from the morphing leading edge structure developed by the LARCASE laboratory at ETS Montreal. The results showed that by means of the optimization strategy adopted, the force required on the actuator mechanism is 88% lower than the original design. A significant improvement on the profile smoothness along its section and in the spanwise direction in morphing conditions has been obtained too. However, further investigations are still needed in order to achieve a more appropriate morphing shape. Despite this, it appears from the results obtained that the proposed methodology can be useful to tackle the DNLE design problem in an effective and efficient way. What developed in this work has been conceived to support the investigation of DNLE in the small leading edge profiles typical of the UAS. In this way, an easier procedure for the set up of the design flow, and a decrease in the computational effort for the optimization process can be obtained. An experimental validation of the results obtained is currently being performed at ETS, and future development regards the assessment of the errors of the numeric procedure herein presented respect to real data.
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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.
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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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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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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.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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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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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.Applied Science, Faculty of
Mechanical Engineering, Department of
Graduate
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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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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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Kaufmann, Markus. "Cost Optimization of Aircraft Structures." Doctoral thesis, KTH, Lättkonstruktioner, 2009. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-11482.
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Composite structures can lower the weight of an airliner significantly. Due to the higher process complexity and the high material cost, however, the low weight often comes with a significant increase in production cost. The application of cost-effective design strategies is one mean to meet this challenge. In this thesis, a simplified form of direct operating cost is suggested as a comparative value that in combination with multidisciplinary optimization enables the evaluation of a design solution in terms of cost and weight. The proposed cost optimization framework takes into account the manufacturing cost, the non-destructive testing cost and the lifetime fuel consumption based on the weight of the aircraft, thus using a simplified version of the direct operating cost as the objective function. The manufacturing cost can be estimated by means of different techniques. For the proposed optimization framework, feature-based parametric cost models prove to be most suitable. Paper A contains a parametric study in which a skin/stringer panel is optimized for a series of cost/weight ratios (weight penalties) and material configurations. The weight penalty (defined as the specific lifetime fuel burn) is dependent on the fuel consumption of the aircraft, the fuel price and the viewpoint of the optimizer. It is concluded that the ideal choice of the design solution is neither low-cost nor low-weight but rather a combination thereof. Paper B proposes the inclusion of non-destructive testing cost in the design process of composite components, and the adjustment of the design strength of each laminate according to inspection parameters. Hence, the scan pitch of the ultrasonic testing is regarded as a variable, representing an index for the guaranteed material quality. It is shown that the cost for non-destructive testing can be lowered if the quality level of the laminate is assigned and adjusted in an early design stage. In Paper C and Paper D the parameters of the manufacturing processes are upgraded during the cost optimization of the component. In Paper C, the framework is extended by the cost-efficient adaptation of parameters in order to reflect the situation when machining an aluminum component. For different weight penalties, the spar thickness and stringer geometry of the provided case study vary. In addition, another cutter is chosen with regard to the modified shape of the stringer. In Paper D, the methodology is extended to the draping of composite fabrics, thus optimizing not only the stacking layup, but also the draping strategy itself. As in the previous cases, the design alters for different settings of the weight penalty. In particular, one can see a distinct change in fiber layup between the minimum weight and the minimum cost solution. Paper E summarizes the work proposed in Papers A-D and provides a case study on a C-spar component. Five material systems are used for this case study and compared in terms of cost and weight. The case study shows the impact of the weight penalty, the material cost and the labor rate on the choice of the material system. For low weight penalties, for example, the aluminum spar is the most cost-effective solution. For high weight penalties, the RTM system is favorable. The paper also discusses shortcomings with the presented methodology and thereby opens up for future method developments.QC 20100723
European Framework Program 6, project ALCAS, AIP4-CT-2003-516092
Nationella flygtekniska forskningsprogrammet (NFFP) 4, project kostnadseffektiv kompositstruktur (KEKS)
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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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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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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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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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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.Includes bibliographical references (leaves 89-91).
by Ian B. Land.
S.M.
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Bendemra, Hamza. "Automation of Lay-Up in the Repair of Advanced Composite Aircraft Structures." Phd thesis, Canberra, ACT : The Australian National University, 2016. http://hdl.handle.net/1885/101713.
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The use of lightweight composite materials in aircraft structure
has become increasingly widespread over the past thirty years as
the need for reduced fuel consumption and improved performance
grew stronger. This has also raised concerns in the event of
damage as damage patterns in composite structures can be
unpredictable and difficult to detect when located under the
surface. The high cost of fibre-reinforced composite materials
has made replacement a less attractive option. Therefore, the
need for an efficient and cost-effective repair method for
composite structures has become significant. Nowadays, most
composite repair operations are undertaken manually by repair
engineers and technicians. However, manual composite repair is
time-consuming and requires extensive training.
The challenge is to automate the process to reduce repair time
and increase efficiency while maintaining strict aerospace
quality requirements. The accuracy and repeatability offered by
an automated process has the potential to meet such requirements.
Previous research efforts have mainly focused on automated
scarfing and automated inspection methods. Further research is
required for the automation of composite repair patch
manufacturing and application. This research project, supported
by Boeing Research and Technology Australia, aims to complement
global research efforts on automated composite repair. Two
research aims were identified. Firstly, determine optimised
repair patch shapes including joint parameters which are suitable
for automated patch manufacturing. Secondly, investigate
manufacturability and feasibility of composite repair patch
manufacturing using the AFP method.
Extensive finite element modelling was performed to determine
optimised repair shapes suitable for AFP. Two optimised repair
shapes were identified: the octagon and the square-ellipse. The
octagonal shape reduced the creation of adhesive rich areas at
the parent-patch interface with composite prepreg tows in fibre
directions used in this study (i.e. 0, ±45, 90). Finite element
analysis was then performed for the optimised repair shapes.
Stress results showed that the optimised shapes provided strength
and stiffness in highly loaded areas while significantly reducing
overall repair size compared to the traditional circular patch.
An experimental AFP apparatus was developed in-house and
subsequently used for AFP composite repair patch manufacturing.
Three-point bending tests were performed to characterise flexural
strength for each repair shape. Experimental results validated
the feasibility of using AFP for repair patch fabrication, and
further strengthen the case for optimised AFP repair
configurations which showed promising flexural strength results,
particularly when compared to the traditional circular repair
patch shape currently used in the industry. Several research
opportunities have emerged from this research project which can
be addressed in subsequent projects. They include the development
of a mobile repair unit; the implementation of AFP repair patch
fabrication for thermoset composite structures; the investigation
of the elastic plastic behaviour of the adhesive in the optimised
repair shapes under hot/wet environmental conditions; and
opportunities for improvements to the in-house AFP equipment.
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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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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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Lazarin, Juan Reuben. "Optimum Design of Composite Wing Spar Subjected to Fatigue Loadings." DigitalCommons@CalPoly, 2017. https://digitalcommons.calpoly.edu/theses/1816.
Full textAbstract:
Composites are now being incorporated into aircraft designs because of their high strength to weight ratio compared to traditional metal materials. Due to the complexity of the material, composite parts are presently being over designed to satisfy static and fatigue requirements. A greater understanding of composite fatigue behavior will allow for even greater weight savings leading to increased fuel economy. A critical part of an aircraft that is subjected to fatigue bending loads are its wings. The forces acting on the wings include its lift distribution, powerplant, and fuel which can be carried in the wing body. When in flight these forces repeatedly cause cyclic displacements which could ultimately lead to failure. It is important to design the wing spars which carry the bending loads, to be fatigue resistant so that damage or expensive inspections could be avoided.
Wing models were be made from composite materials with a NACA 0016 airfoil shape, chord length of 9.25”, and a span of 15.25”. The C – channel spars were located at 22% and 72% of the chord. Strain gages on the wing model were used to measure strain at different locations. Static test were conducted on the specimens in order to validate a finite element analysis(FEA) model to be used for simulations. Overall, the strain measurements on the leading edge from two of the wings matched the model within 9% of the simulation results. Additional spar designs were then analyzed to determine the optimal one for static and fatigue bending loads. The wings were fatigue tested under displacement control at a test frequency. A model 8801 servo-hydraulic Instron machine and Wave Matrix software was used to fatigue the wings. After 100,000 cycles the test would be deemed a success and concluded.
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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/.
Full textAbstract:
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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Mohammed, Bizuayehu Y. "Damage characterisation in twill-weave CFRP composite aircraft structures using modal analysis of acoustic emission signals." Thesis, University of South Wales, 2016. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.702330.
Full textAbstract:
The aim of this research work is to propose a damage characterisation method using acoustic emission technology. The research is broadly categorised in structural health monitoring (SHM) of CFRP composite structures. A number of tests were performed on twill-weave CFRP composite material in correlating the various damage types to their associated signal features. Recent developments in aircraft structural material lead to an increase use of CFRP composite structures. However, these materials are susceptible to different modes of failures such as matrix cracking, delamination, stringer debonding, and fibre fracture which affect the integrity of the structure. Determining the type of failure during service life of an aircraft structure is an important input in the SHM of materials. This can potentially reduce inspection time, and increases knowledge of the damage type propagating within the said structure. Acoustic emission (AE) is a phenomenon where a stress wave is generated due to stresses in a material. AE sensors can be placed on the surface of the structure to detect these waves while the material is in service. These waves have possess distinct signal features which can be attributed to a particular damage type. Therefore, AE based technology is potentially suitable for the SHM of composite structures. This thesis proposes an improved damage characterisation method for twill weave CFRP composite when subjected to various modes of failure. These failure modes were achieved using a variety of test setups namely cantilever bending, three point bending, stringer debonding, and tensile testing. The proposed method uses live digital microscope video recording of events along with AE events. Distinct signal features were identified for delamination initiation and propagation, matrix cracking, fibre fracture and skin-stringer debonding.
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Seitz, Timothy J. "Formulation of a structural model for flutter analysis of low aspect ratio composite aircraft wings." Diss., This resource online, 1992. http://scholar.lib.vt.edu/theses/available/etd-05042006-164511/.
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Matt, Howard M. "Structural diagnostics of CFRP composite aircraft components by ultrasonic guided waves and built-in piezoelectric transducers." Connect to a 24 p. preview or request complete full text in PDF format. Access restricted to UC campuses, 2007. http://wwwlib.umi.com/cr/ucsd/fullcit?p3238426.
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Thesis (Ph. D.)--University of California, San Diego, 2007.Title from first page of PDF file (viewed January 4, 2007). Available via ProQuest Digital Dissertations. Vita. Includes bibliographical references (p. 203-216).
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Clermont, Paul Daniel Stanley. "Characterization and prediction of flow electrification phenomena in fuel tanks of aeronautical structures." Thesis, Poitiers, 2016. http://www.theses.fr/2016POIT2258/document.
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Avec la nouvelle génération d'avions composites, une attention est portée sur les systèmes de carburant vis-à-vis de la prévention des décharges électrostatiques (ESD) durant les phases de remplissage des réservoirs. La plupart des travaux réalisés en aéronautique a été menée sur des réservoirs métalliques. Toutefois, l'introduction des matériaux composites a soulevé de nouvelles interrogations, puisque ces matériaux peuvent avoir un comportement différent des métaux vis-à-vis de l'électrisation par écoulement, qui justifient pleinement de nouvelles analyses. Afin de définir correctement les structures des réservoirs et leur protection contre les risques ESD, il est crucial de comprendre comment un empilement complexe de matériaux se comporte en termes de création de charge lorsque ces matières sont en contact avec un carburant d'avion. La structure de ces matériaux et leurs propriétés électriques contrôlent le potentiel électrique atteint dans le réservoir à travers un équilibre entre la production, l'accumulation et la fuite des charges électriques. Ce potentiel peut dépasser le point d'éclair du mélange air/vapeurs de carburant et provoquer une inflammation. Diverses mesures de protection peuvent être adoptées pour contrôler ce phénomène, comme utiliser des additifs antistatiques dans les carburants, des réservoirs métalliques à la masse ou encore des réservoirs faits de matériaux non métalliques mais ne favorisant pas l'accumulation de charges. C'est principalement en réponse à cette dernière solution que ce travail est orienté afin de guider vers le choix optimaux des matériaux et une meilleure définition des structures du réservoirWith the new generation of composite aircrafts an attention is carried out on fuel systems with respect to prevention of electrostatic discharges (ESD) during the filling phases of the tanks. Most of the work realized in aeronautics (during the 60's) was conducted on metallic fuel tanks. However, the introduction of composite materials has raised new questions, since those materials can have a different behavior than metallic ones with respect to flow electrification, which fully justify new analyses. In order to properly define the tank structures and their protection against ESD hazards, it is crucial to understand how a complex stack of materials (conductive or not, multilayered or homogeneous, painted or not) constituting a fuel tank behaves with respect to the mechanisms of charge creation by flow electrification when these materials are in contact with aviation fuel. The structure of these materials and their electrical properties control the electric potential reached in the tank through a balance between the production, accumulation and leakage of the electrical charge. This potential may exceed the flash point of the fuel vapors/air mixture and induce ignition. Various protective measures can be adopted to control this phenomenon such as using antistatic additives in the fuels, lowering the rates ofthe fuel injection inside the tank, using only bonded metallic tanks or tanks made of non-metallic materials which do not favor charge accumulation or local charge trapping. It is majorly in response to the latter solution that this work is oriented in order to guide optimum choices of materials and a better definition of the tank structures
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Isbilir, Ozden. "Optimisation of the geometry of the drill bit and process parameters for cutting hybrid composite/metal structures in new aircrafts." Thesis, University of Sheffield, 2012. http://etheses.whiterose.ac.uk/2890/.
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Vandeveld, Thierry F. R. "Etude expérimentale multisensorielle de la dynamique des impacts d'oiseaux sur structures d'avions." Doctoral thesis, Universite Libre de Bruxelles, 2009. http://hdl.handle.net/2013/ULB-DIPOT:oai:dipot.ulb.ac.be:2013/210274.
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Chaque année, d'innombrables collisions se produisent entre des avions en vol et des oiseaux. L'impact aviaire, menace redoutée par les pilotes, concerne tant l'aviation civile que son pendant militaire. Les statistiques démontrent que, même si fort heureusement le nombre d'accidents graves reste limité, les incidents sont de plus en plus nombreux.Parmi les acteurs qui luttent contre ce danger, les constructeurs d'avions jouent un rôle prépondérant. Contraints par des réglementations internationales, ils s'attachent à produire des éléments de structure qui résistent à l'impact d'oiseaux.
Dans la mise au point de leur produits, les avionneurs démontrent cette résistance à l'aide d'essais d'impact :on accélère un simulant d'oiseau jusqu'à la vitesse voulue -- de l'ordre de la vitesse de croisière nominale de l'avion -- et on le projette sur un aileron ou un morceau de fuselage.
La présente thèse doctorale, co-dirigée par les professeurs Philippe Bouillard de l'ULB et Marc Pirlot de l'ERM, contribue doublement à l'amélioration de ces techniques d'essais dynamiques.
D'une part, elle réalise la mise au point et la validation d'un lanceur pyrotechnique à double étage pour l'accélération du simulant d'oiseau. Un canon de calibre 20 mm est combiné avec un accélérateur de calibre 160 mm. La combustion d'un mélange de poudre propulsive contenu dans une douille adaptée génère les gaz à haute température et à haute pression nécessaires à l'accélération d'un simulant d'oiseau dûment confiné dans un conteneur de protection. Un dispositif de séparation arrête le conteneur afin que seul le simulant d'oiseau percute l'élément d'avion à l'essai. La solution pyrotechnique à double étage mise au point est validée par de nombreux tirs instrumentés en vitesse, en accélération et en pression ;elle se révèle conforme aux exigences de sécurité et de reproductibilité. Le lanceur pyrotechnique présente par rapport aux solutions pneumatiques, utilisées à notre connaissance dans tous les autres centres d'essais, des avantages indéniables de compacité ainsi que de rapidité et de souplesse de mise en oeuvre.
D'autre part, la migration des alliages métalliques vers les matériaux composites est amorcée depuis plusieurs années déjà dans le monde de la construction aéronautique. Pour optimiser les structures, une connaissance des caractéristiques de ces matériaux est indispensable. Les modes de rupture font partie des caractéristiques encore mal connues. La mesure du déplacement hors-plan lors du tir sur panneaux plans est une des manières de quantifier le comportement du matériau sous l'action d'un impact. Cette mesure s'opère généralement de manière statique, après le tir. Une méthode de mesure dynamique a été mise au point, basée sur l'emploi de techniques de stéréoscopie par corrélation numérique d'images. Cette technique a été validée au moyen d'une méthode métrologique indépendante d'extensométrie laser.
ABSTRACT
Countless collisions occur each year between airplanes and birds. Bird strike is a concern to both civilian and militay aircraft. Statistics show that, although the number of serious accidents fortunately remains low, the number of incidents keeps increasing.
Amongst the actors tackling this issue, aircraft manufacturers play an important role. In compliance with international regulations, they have to produce structural elements that withstand bird impact. During the development of their products, aircraft manufacturers have to demonstrate this resistance through bird impact trials :a bird surrogate is accelerated to the required velocity - often close to the nominal cruise speed of the aircraft - and launched onto a flap or a piece of fuselage.
This PhD thesis has been co-supervised by Professor Philippe Bouillard (ULB) and Professor Marc Pirlot (ERM-KMS). Its contribution to the improvement of the aforementioned dynamic trials is twofold.
One one hand, a two-stage pyrotechnical launcher for bird surrogates has been developed and assessed. A 20 mm caliber gun is connected to a 160 mm diameter launcher. The combustion of a propellant mixture in a cartridge case generates high pressure, high temperature gases which accelerate a bird surrogate protected by a cylindrical container. A stripper refrains the container from hitting the target pane.
The pyrotechnical solution has been assessed through an important number of firings where pressure, velocity and acceleration have been measured. The solution has proven compliance with both the safety requirements and the repeatability specifications. Its advantages compared to the pneumatic solutions used, as far as we know, in all other test centres, include compactedness as well as flexibility and high firing rate.
On the other hand, migration towards composite materials has been initiated years ago in the area of aeronautical constructions.
To optimize structures, a thorough knowledge of these new materials is required. Failure modes belong to the still badly known features of carbon reinforced plastics. Measuring the out-of-plane deformation when firing on a flat pane is one way of quantifying the material's behaviour under impact. This measurement is most frequently made in a static way, after completion of the firing. A dynamic measuring method has been developed, based upon stereoscopic digital image correlation techniques. This technique has been validated by means of an independent laser extensometer measuring method.
Doctorat en Sciences de l'ingénieur
info:eu-repo/semantics/nonPublished
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Sakarya, Arzu. "Multidisciplinary Design Of An Unmanned Aerial Vehicle Wing." Master's thesis, METU, 2011. http://etd.lib.metu.edu.tr/upload/12613606/index.pdf.
Full textAbstract:
In this thesis, the structural design, structural analysis and producibility analysis of an unmanned aerial vehicle wing were performed. Three different wing models, made of different materials, were designed. The wings were aluminum wing model and composite wing modelsmade of prepreg and wet lay-up. All wings have the same aerodynamic geometry and structural configuration under the same flight conditions. The structural designs of three wings were done by using Unigraphics NX. The finite element modeling of the wings were built by using MSC Patran package program. After the application of the loads on models, structural analyses were performed by MSC Nastran. Finally, the producibility analysis of prepreg wing model was conducted by using FiberSIM package program. The prepreg wing model was selected as optimum design with studies conducted in the study considering weight, producibility, cruise and gust stress and displacement conditions.
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Foschini, Lorenzo. "Studio della produzione di provini laminati manualmente per prove di caratterizzazione meccanica." Bachelor's thesis, Alma Mater Studiorum - Università di Bologna, 2019.
Find full textAbstract:
I materiali compositi nascono dall'unione di due o più materiali differenti, che conferisce al prodotto finale proprietà superiori rispetto a quelle dei singoli.
Grazie alle qualità dimostrate e all’intercambiabilità degli elementi costituenti, è stato possibile allargare il campo di utilizzo dei materiali compositi alla maggior parte dei settori dell’industria moderna; uno dei settori che maggiormente risente dell’influenza di questi materiali è il settore aereonautico, che, grazie all’implementazione di tecnologie, è riuscito ridurre drasticamente la presenza dei materiali convenzionali, ottimizzando le caratteristiche del velivolo stesso.
In questo elaborato sono stati presi in considerazione i CFRP (carbon/epoxy) in configurazione cross-ply, approfondendone le procedure di corretta laminazione tramite una campagna sperimentale.
L’obiettivo è stato produrre una serie di provini da sottoporre a verifiche di resistenza meccanica a compressione, sia in condizioni ottimali, sia simulando condizioni di service per mezzo di una serie d’impatti a basse velocità (BVID – Barely Visible Impact Damages).
La laminazione e le procedure d’impatto si sono svolte presso il laboratorio hangar dell’università di Ingegneria e Architettura con sede a Forlì, nel quale è stato possibile effettuare un processo di laminazione in autoclave e compiere impatti centrali al provino e nelle prossimità del bordo.
La caratterizzazione del materiale, sottoposto a prove di compressione, è stata effettuato tramite l’impiego di attrezzature all’avanguardia presso il centro di ricerca ENEA, Unità Tecnica Tecnologie dei Materiali di Faenza.
Il seguente elaborato descrive i mezzi e le considerazioni necessarie per portare a termine l’attività sperimentale.
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Wu, Ching-Shiuh, and 吳清旭. "Use of embedded optical fibers for damage detection in aircraft composite structure." Thesis, 1993. http://ndltd.ncl.edu.tw/handle/03685677949644815214.
Full textAbstract:
碩士國立臺灣大學
材料科學(工程)研究所
81
Optical fibers were embedded in the graphite/ epoxy (Gr/Ep) laminates to study their possibility of use as a sensor to detect the structural damage of the laminates. To improve their reliability. Optical fibers were surface-treated by hydrogen fluoride (HF) and/or silane solutions. According to the Weibull statistical analysis and electronic microscopy, as the optical fibers were treated by HF solution,their tensile failure was changed from intrinsic failure to extrinsic failure so that the distribution of tensile strength was narrower. As to the adhesion of optical fibers to the epoxy resins, the one treated with HF solution for 10 minutes and then coated with silane showed the highest adhesion strength. During the lamination , a laser beam was injected to an embedded optical fiber, and found that the light in tensity was decreased with increasing temperature and pressure. However, after the laminate was cured and returned to the ambient environment, the light intensity could be recovered to the original intensity, indicating that the optical fiber did not damage during lamination. Finally, to detect the structural damage of optical fiber-embedded laminates during tensile testing at an extremely slow rate, a transmitted light intensity was recorded. Experimental results revealed that if the optical fibers were treated with HF solution and then coated with silane they were more sensitive to detect the damage of lami- nates.
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王怡安. "Fabrication of Stitched Braided Composite Aircraft Window Frame Structure by Resin Film Infusion." Thesis, 2013. http://ndltd.ncl.edu.tw/handle/zc6drv.
Full textAbstract:
碩士逢甲大學
纖維與複合材料學系
102
The aim of this project is to develop a manufacturing technology of high performance aircraft window frame composite structure and discuses the interlaminar shear strength and fracture toughness of structure. Combination of stitch-braided and resin film infusion (RFI) processes to fabrication carbon stitch-braided aircraft window frame composite structure. Whole the project, including stitching and braiding and resin film infusion techniques were studied and detail investigate the interlaminar shear strength (ILSS) and fractographic observations of the failure of stitch-braided composites and the fabrication of aircraft window frame composite structure as well studied. The effects of stitching factor and process parameters on the properties of composite structures were studied. Resin film infusion (RFI) technology for manufacturing aircraft window frame composite structures and the effects of stitching parameters, such as; stitching yarn type, stitching density, stitching layer, stitching space and stitching tension on the stitching process were studied. The fracture toughness (GI and GII) of stitch-braided composites tested by mixed fracture modes test and ILSS were used to evaluate the stitched and braiding process conditions. As well as, the torsion property and compression property and fractographic observations of the failure of window frame composite structure were discussed. Finally, the manufacturing processes technologies of aircraft window frame composite structures were established
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Κοτζακόλιος, Αθανάσιος. "Blast response of aircraft structures." Thesis, 2011. http://hdl.handle.net/10889/5068.
Full textAbstract:
The scope of this project is the realization of composite and hybrid sub-aerostructures which exhibit superior blast performance compared to reference composite and hybrid substructures. The scope will be fulfilled with minimum weight penalty. Within the scope of this work is to provide a roadmap for the integration of explicit hardening measures for blast in future aerospace structural components. In the case of blast loading, the proposed methodology for achieving these aims involves vulnerability analysis of the composite and the hybrid substructures (scaled fuselage substructure). The vulnerability analysis will be based on numerical results, obtained by the systematic, analysis of the coupled blast / structural problem. The aims and objectives of the present project can be summarized as follows:
• Development of numerical models and their correlation against experimental results.
• Development of numerical tools for blast vulnerability analysis of composite and hybrid aeronautic structures
• Blast vulnerability map of composite and hybrid scaled fuselage substructure for different charge locations
• Explicit blast hardening strategies of composite and hybrid aerostructures by design and by novel materialsΣκοπός της εργασίας αυτής είναι η μελέτη αεροπορικών κατασκευών από σύνθετα υλικά όπου θα παρουσιάζουν βελτιωμένες ιδιότητες υπό συνθήκες έκρηξης σε σύγκριση με υπάρχουσες αντίστοιχες κατασκευές.Ο στόχος αυτός επετεύχθη με ελάχιστη προσθήκη βάρους. Μέσα στους στόχους της παρούσας διατριβής είναι να παρέχει μια μεθοδολογία για την προσθήκη μέσων προστασίας για τις μελλοντικές αεροπορικές κατασκευές.
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Ajith, V. "Wave Propagation in Healthy and Defective Composite Structures under Deterministic and Non-Deterministic Framework." Thesis, 2012. http://hdl.handle.net/2005/3253.
Full textAbstract:
Composite structures provide opportunities for weight reduction, material tailoring and integrating control surfaces with embedded transducers, which are not possible in conventional metallic structures. As a result there is a substantial increase in the use of composite materials in aerospace and other major industries, which has necessitated the need for structural health monitoring(SHM) of aerospace structures. In the context of SHM of aircraft structures, there are many areas, which are still not explored and need deep investigation. Among these, one of the major areas is the development of efficient damage models for complex composite structures, like stiffened structures, box-type structures, which are the building blocks of an aircraft wing structure. Quantification of the defect due to porosity and especially the methods for identifying the porous regions in a composite structure is another such area, which demands extensive research. In aircraft structures, it is not advisable for the structures, to have high porosity content, since it can initiate common defects in composites such as, delamination, matrix cracks etc.. In fact, there is need for a high frequency analysis to detect defects in such complex structures and also to detect damages, where the change in the stiffness due to the damage is very small. Lamb wave propagation based method is one of the efficient high frequency wave based method for damage detection and are extensively used for detecting small damages, which is essentially needed in aircraft industry. However, in order, to develop an efficient Lamb wave based SHM system, we also need an efficient computational wave propagation model. Developing an efficient computational wave propagation model for complex structures is still a challenging area. One of the major difficulty is its computational expense, when the analysis is performed using conventional FEM. However, for 1D And 2D composite structures, frequency domain spectral finite element method (SFEM), which are very effective in sensing small stiffness changes due to a defect in a structure, is one of the efficient tool for developing computationally efficient and accurate wave based damage models. In this work, we extend the efficiency of SFEM in developing damage models, for detecting damages in built-up composite structures and porous composite structure. Finally, in reality, the nature of variability of the material properties in a composite structure, created a variety of structural problems, in which the uncertainties in different parameters play a major part. Uncertainties can be due to the lack of good knowledge of material properties or due to the change in the load and support condition with the change in environmental variables such as temperature, humidity and pressure. The modeling technique is also one of the major sources of uncertainty, in the analysis of composites. In fact, when the variations are large, we can find in the literatures available that the probabilistic models are advantageous than the deterministic ones. Further, without performing a proper uncertain wave propagation analysis, to characterize the effect of uncertainty in different parameters, it is difficult to maintain the reliability of the results predicted by SFEM based damage models. Hence, in this work, we also study the effect of uncertainty in different structural parameters on the performance of the damage models, based on the models developed in the present work.
First, two SFEM based models, one based on the method of assembling 2D spectral elements and the other based on the concept of coupling 2D and 1D spectral elements, are developed to perform high frequency wave propagation analysis of some of the commonly used built-up composite structures. The SFEM model developed using the plate-beam coupling approach is then used to model wave propagation in a multiple stiffened structure and also to model the stiffened structures with different cross sections such as T-section, I-section and hat section.
Next, the wave propagation in a porous laminated composite beam is modeled using SFEM, based on the modified rule of mixture approach. Here, the material properties of the composite is obtained from the modified rule of mixture model, which are then used in SFEM to develop a new model for solving wave propagation problems in porous laminated composite beam. The influence of the porosity content on the parameters such as wave number, group speed and also the effect of variation in theses parameters on the time responses are studied first. Next, the effect of the length of the porous region (in the propagation direction) and the frequency of loading, on the time responses, is studied. The change in the time responses with the change in the porosity of the structure is used as a parameter to find the porosity content in a composite beam.
The SFEM models developed in this study is then used in the context of wave based damage detection, in the next study. First ,the actual measured response from a structure and the numerically obtained response from a SFEM model for porous laminated composite beam are used for the estimation of porosity, by solving a nonlinear optimization problem. The damage force indicator (DFI) technique is used to locate the porous region in a beam and also to find its length, using the measured wave propagation responses. DFI is derived from the dynamic stiffness matrix of the healthy structure along with the nodal displacements of the damaged structure. Next, a wave propagation based method is developed for modeling damage in stiffened composite structures, using SFEM, to locate and quantify the damage due to a crack and skin-stiffener debonding. The method of wave scattering and DFI technique are used to quantify the damage in the stiffened structure.
In the uncertain wave propagation analysis, a study on the uncertainty in material parameters on the wave propagation responses in a healthy metallic beam structure is performed first. Both modulus of elasticity and density are considered uncertain and the analysis is performed using Monte-Carlo simulation (MCS) under the environment of SFEM. The randomness in the material properties are characterized by three different distributions namely normal, Weibul and extreme value distribution and their effect on wave propagation, in beam is investigated. Even a study is performed on the usage of different beam theories and their uncertain responses due to dynamic impulse load.
A study is also conducted to analyze the wave propagation response In a composite structure in an uncertain environment using Neumann expansion blended with Monte-Carlo simulation (NE-MCS) under the environment of SFEM. Neumann expansion method accelerates the MCS, which is required for composites as there are many number of uncertain variables. The effect of the parameters like, fiber orientation, lay-up sequence, number of layers and the layer thickness on the uncertain responses due to dynamic impulse load, is thoroughly analyzed.
Finally, a probabilistic sensitivity analysis is performed to estimate the sensitivity of uncertain material and fabrication parameters, on the SFEM based damage models for a porous laminated composite beam. MCS is coupled with SFEM, for the uncertain wave propagation analysis and the Kullback-Leibler relative entropy is used as the measure of sensitivity. The sensitivity of different input variables on the wave number, group speed and the values of DFI, are mainly considered in this study.
The thesis, written in nine chapters, presents a unified document on wave propagation in healthy and defective composite structure subjected to both deterministic and highly uncertain environment.
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Clark, Randal John. "Prediction of fatigue damage progression in bonded composite repairs to aluminum aircraft structures." Thesis, 2000. http://hdl.handle.net/2429/10395.
Full textAbstract:
The focus of this thesis is the development of a predictive model for fatigue damage progression
in unidirectional bonded composite repairs of cracked isotropic plates. The principal use of this
technology is in the design of repairs for aluminum aircraft structures. The ability to predict the
rate of fatigue damage is critical to damage tolerance analysis of a repair. A damage tolerance
analysis will allow designers to assess the design life, assign inspection intervals, and determine
likely failure modes for a repair, and will be required for airworthiness certification for a long or
indefinite period of operation.
In this thesis, classical methods of bonded joint analysis are presented, and extended to the case
of reversed plasticity of an internally pressurized lap joint. The crack bridging effect of the repair
is examined using a boundary element method employing the Green's functions for a point load
applied to a center-cracked plate. This results in a system of linear equations solvable by Gauss-
Seidell iteration. The boundary element method allows calculation of the stress intensity and
adhesive shear stresses under the bonded patch. These parameters govern the fatigue and static
strength of the repair. The boundary element model combines engineering fracture mechanics
and bonded joint analysis techniques in a very direct and straightforward manner.
Results from the boundary-element model are compared to approximate analytical methods for a
disbonding patch, and an improved analytical model employing correction factors is presented.
Power law methods are then used to predict crack and disbond growth rates, which are compared
to experimental results. The influence on patch life of various secondary effects, such as adhesive
plasticity, process-induced thermal residual stresses, patch bending, shear deformation of the
patch, and cracked-plate geometry are investigated.
Based on this work, conclusions are drawn regarding patch behavior, limitations of modeling
techniques, and experimental results still necessary to validate patch mechanics models. The
techniques developed are also of interest in the study of cracking in fiber-reinforced metal
laminates (FRML).
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Robertson, Cameron David. "Structural Characterization, Optimization, and Failure Analysis of a Human-powered Ornithopter." Thesis, 2009. http://hdl.handle.net/1807/18847.
Full textAbstract:
The objective of this work was to develop an analysis framework for the structural design
of the Human-Powered Ornithopter (HPO). This framework was used in a kinematicaerostructural
optimizer for
apping-wing
ight (Ornithia), as well as analytically to design
the HPO, and focused on three goals. First was the development of an accurate and computationally
inexpensive nite-element method, to be integrated with Ornithia, which would
capture the geometric nonlinearity of the aerostructural interaction of the wing when subjected
the large deformations in
ight. Second was the assembly of a model by which the
aircraft primary structure, the wing main spar especially, could be exactly characterized and
designed. Third was the establishment of a process and toolbox for failure analysis which
could be applied universally in the design of the HPO. The validation and tuning of these
models involved extensive testing on prototype carbon ber composite components.
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Salva, Joao Manuel Correia. "A load/cost/mass comparison of aluminium, glass, carbon and asbestos fibre composites for immediate application to aircraft primary structures." Thesis, 2015. http://hdl.handle.net/10539/18603.
Full textAbstract:
A Dissertation Submitted to the Faculty of Engineering,
University of the W i twatersrand, Johannesburg in fulfilment
of the requirements for the degree of Master
of Science in Engineering.
Johannesburg 1979.A comparison of 4 different materials for immediate app l i cation to aircraft primary structures was made on the basis of load-to-buckle/cost/mass. The materials tested were: Al-2024 T3 (Alclad) , glass fibre cloth, carbon fibre and asbestos fibre (Noramite) composites. A m a t h e matical formulation of the problem was used which was found to give very satisfactory res ults. This method made use of beam vibration modes and the Rayleigh-Ritz energy formulation and it was found that even with onl\ 4 modes of vibration, the results agreed very well with the computer analysis. A working equation lor the design of composite panels in shear is also given.
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Sudha, J. "Fatigue Damage Characterization Of Carbon/Epoxy Laminates Under Spectrum Loading." Thesis, 2007. http://etd.iisc.ernet.in/handle/2005/2257.
Full textAbstract:
Fibre Reinforced Polymer Composites are extensively used in aircraft structures because of its high specific stiffness, high specific strength and tailorability. Though Fibre Reinforced Polymers offer many advantages, they are not free from problems. The damage of different nature, e.g., service mechanical damages, fatigue damage or environmental damage can be observed during operating conditions. Among all the damages, manufacturing or service induced, delamination related damage is the most important failure mechanisms of aircraft-composite structures and can be detrimental for safety. Delamination growth under fatigue loading may take place due to local buckling, growth from free edges and notches such as holes, growth from ply-drops and impact damaged composites containing considerable delamination. Delamination growth can also occur due to interlaminar stresses, which can arise in complex structures due to unanticipated loading.
The complex nature of composite failure, involving different failure modes and their interactions, makes it necessary to characterize/identify the relevant parameters for fatigue damage resistance, accumulation and life prediction. An effort has been made in this thesis to understand the fatigue behavior of carbon fibre reinforced epoxy laminates under aircraft wing service loading conditions. The study was made on laminates with different lay-up sequences (quasi-isotropic and fibre dominated) and different geometries (plain specimen, specimen with a hole and ply-drop specimen).
The fatigue behaviour of the composite was analyzed by following methods:
. Ultrasonic C-Scan was used to characterize the delamination growth.
. Dynamic Mechanical Analysis (DMA) was done to study the interfacial degradation due to fatigue loading. In this analysis, the interfacial strength indicator and interfacial damping were calculated. The DMA also provides the storage modulus degradation under fatigue loading.
. Scanning electron microscope examination was carried out to understand the fatigue damage mechanisms.
. A semi-empirical phenomenological model was also used to estimate the residual fatigue life.
This research work reveals that the Carbon Fibre Reinforced Polymer laminates are in the safe limit under service loading conditions, except the specimen with a hole. The specimen with a hole showed delaminations around the hole due to stress concentration and higher interlaminar stresses at the hole edges and this delamination is found to be associated with fibre breakage and fibre pullout. The quasi-isotropic laminate is found to show poorer fatigue behaviour when compared to fibre dominated laminate and ply-drop also shows poor performance due to high stress concentration in the ply-drop region.
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Duport, Nicolas. "Étude de la stabilité dimensionnelle d'un revêtement polymère sur placage de bois pour structures intérieures d'avions." Thèse, 2016. http://hdl.handle.net/1866/19906.
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