1,720,999 research outputs found
Optical Coherence Tomography for Material Characterization
No full textOptical coherence tomography (OCT) is a non-invasive, contactless and high resolution imaging method, which allows the reconstruction of two or three dimensional depth-resolved images in turbid media. In the past 20 years, OCT has been extensively developed in the field of biomedical diagnostics, while OCT in the non-destructive testing (NDT) field is lagging far behind. The aim of this thesis is to use OCT as a novel NDT technique for material structure characterization and damage detection. Besides an overview of the OCT fundamentals and developments, the thesis is mainly driven by three tasks: instrument design, signal processing, and applications. An integrated OCT system working at a wavelength of 1550 nm has been built. It combines time domain (TD) and Fourier domain (FD) OCT to make the both types of measurement available in one instrument. TD-OCT has the advantage of a large A-scan range and high SNR, while FD-OCT system has tremendous superiority in fast imaging due to no movement of the reference mirror. These two kinds of measurement can be implemented based on the user request in the developed hybrid OCT system, e.g. improved imaging depth or speed. In TD-OCT, the envelope detector was selected as the ideal method for the demodulation of each axial signal. A bandpass filter and 2D median filter are applied before and after demodulation, respectively, to reduce OCT system and speckle noise. In FD-OCT, raw data was first processed to remove the influence from the optical source and dark noise of the CCD detectors. It was then linearly resampled to convert to evenly spaced intervals of wavenumber, instead of wavelength. With an inverse Fourier transform, one depth profile was recovered and a cross-sectional image was constructed by accumulating a series of depth profiles. The quality of cross-sectional images can be further improved by merging multiple images with different pathlength offsets. The application of the designed OCT system was mainly focused on glass fiber composites and the microstructure of the specimens was displayed by either cross-sectional or volumetric images. Special attention was given to delamination growth in a glass fiber composite for wind turbine blade applications. The glass fiber composite was tested by incremental loading. Volumetric images obtained by OCT were further processed to reconstruct 3D crack surface profiles, from which a full field view of the delamination crack was given, providing substantial information for the study of crack growth in the composites. Additionally, the study explored the use of optical coherence elastography (OCE) for the deformation measurement of glass fiber composites, for the first time to the best of our knowledge. The developed OCE system based on speckle tracking was first evaluated by a test of ridge body translation. Then experiments were implemented for a set of glass fiber composites under tensile testing and three point bending. The results show that OCE can measure the internal displacements of a glass fiber composite in the range from a few micrometers to hundreds of micrometers. Besides, other applications are also presented in the thesis. These include defects and thickness measurement of polymer coating and the microstructure characterization of a wooden-panel painting. The results show the designed OCT system also has high potential for these alternative applications. Recommendations for further improvement to the OCT design and the applications are presented at the end of the thesis.Aerospace Structures and MaterialsAerospace Engineerin
Piezoelectric Composite Actuators: Modelling of the Static and Dynamic Behaviour
No full textSmart actuators, made of smart materials, are becoming more attractive in many applications because smart materials are not subjected to wear and does not require lubrication during services. Piezoelectric materials are a group of the many attractive smart materials that are being investigated for many applications today. Piezoelectric materials show fast responses, high efficiency/accuracy and operate on a large bandwidth. Composite materials are of interest because of their design flexibility and because they are lighter than other materials commonly used in aircraft and other applications. For the research reported here, a piezoelectric material was embedded in a composite material to form a piezoelectric composite actuator. This research was conducted to expand our knowledge of piezoelectric composite material actuators, and originated from the need to control air flow separation over an airfoil. There is a need to build a profound knowledge about such actuators before they can be implemented in an airfoil, and to understand which parameters influence the behaviour of piezoelectric composite material actuators under static and dynamic operating conditions. The actuators were manufactured and tested under static and dynamic conditions and the experimental results were compared to the finite element models. The models were incorporated with piezoelectric material properties that were determined experimentally. The results showed that a piezoelectric material’s response to an input stimulant, e.g. electric voltage and frequency, influences the actuator’s behaviour.Aerospace Structures and Materials: Structural Integrity & CompositesAerospace Engineerin
The Residual Strength Failure Sequence in Fibre Metal Laminates
No full textThe concept of damage tolerance is a key aspect in ensuring and maintaining safety of an airframe structure over its design life. Developments in materials and structural design have both contributed to improvements in the damage tolerance of modern aircraft structures. Indeed, new developments in metal alloys, composite materials, and hybrid materials such as the Fibre Metal Laminates (FMLs) have all resulted in structures less sensitive to damage and capable to withstand more severe loading conditions. Among other materials, FMLs represent a clear example of damage tolerant hybrid materials, made by bonding thin metal sheets together with fibres embedded in epoxy. Exploiting the damage tolerance capability of FMLs is strictly related to the ability to firstly understand the occurring failure mechanisms, and secondly to be able to accurately describe those mechanisms. In this light, the present dissertation describes the investigation on the residual strength failure sequence in FMLs, and presents the development of an accurate analytical prediction method. The failure sequence is studied in particular for standard Glare laminates, which are relevant laminates for applications in aircraft pressurized fuselages. The developed analytical method has been implemented into two numerical models, considering both through-the-thickness crack and fatigue crack configurations. The developed models are validated against a large number of experimental data, which are also presented in this thesis. The main concept in this dissertation is that the crack growth process in the metallic layers of an FML can be described with the Crack Tip Opening Angle concept (CTOA). This approach includes the contribution of the fibre layers (e.g. fibre failure and fibre bridging) and the associated quasi-static delamination growth. An introduction to FMLs and to all various Glare grades, lay-ups, and manufacturing processes is provided in chapter 2. Some current and future applications for aircraft structures are also discussed in that chapter. A qualitative description of the principal failure mechanisms occurring during the residual strength failure sequence is presented in chapter 3. Based on experimental observations, the metal crack growth mechanisms, permanent plastic deformation, fibre failure and static delamination growth are discussed. All these mechanisms are related to each other, and all contribute to the residual strength of the laminate. The development of the prediction models aimed to be a step forward with respect to previous relevant prediction models available in literature. Therefore, both empirical and analytical prediction models available in literature are presented and discussed in chapter 4. A critical evaluation of those models has pointed out their limitations in applicability and versatility towards a “generic FML” concept. From this chapter, some guidelines have been defined to address the subsequent model development. Two types of experimental activities were carried out. The first type consisted in experiments to gain understanding of the deformation behaviour of both metallic and fibre layers. Extensive use of Digital Image Correlation technique enabled to observe and measure the deformation field of both metal and fibre layers, and their interaction. Further insight into the fibre bridging mechanism and into the metal-fibre interaction was obtained. These experimental activities are discussed in chapter 5. The second type of experimental activities aimed to generate input data for the prediction model, and to validate the CTOA approach. These are discussed in chapter 6. A large amount of experimental CTOA tests were conducted on several FML grades to evaluate the CTOA as failure criterion for FML. This included the investigation of the effect of metal sheet thickness, crack length-to-panel width ratio and the effect of bridging fibres. Static delamination growth tests were conducted to obtain the critical Strain Energy Release Rate. This parameter was subsequently used as input in the prediction model to define the critical condition for the delamination growth. Furthermore, in the same chapter, it is also discussed the complex interaction between static delamination growth and plastic deformation of the metallic layers. The core aspect of the present thesis concerns the modelling of the residual strength failure sequence, which is presented in chapter 7. Two models are described: one for the through-the-thickness crack and one for the fatigue crack. Both models are based on the same method, which uses the CTOA as crack growth driving parameter. The method is based on the idea that crack extension in the metallic layers occurs when the calculated CTOA reaches the critical value obtained from CTOA experiments on metal laminates containing the same metal layers used in the FML. The calculated CTOA is a function of the contribution due to the far-field stress in the aluminium layers, and the contribution of the fibres. The fibre can contribute either in terms of crack opening contribution (broken fibres) or crack closing contribution (bridging fibres present in the fatigue crack configuration). Plastic deformation ahead and behind the metal crack tip is accounted and implemented into the calculation. In addition, in the case of fatigue crack configurations, the bridging stress is calculated by solving the deformation compatibility equation, accounting for the plastic zone ahead of the crack tip and fibre failure in the bridging area. The bridging stress is subsequently used to calculate the quasi-static delamination growth occurring at the fibre-metal interface using the Strain Energy Release Rate approach. The model for through-the-thickness crack showed a very good agreement with the experimental data, while the model for fatigue crack configuration showed sufficient agreement with experimental data. The modelling of the fatigue crack configuration presents higher degree of complexity, which required a number of simplifications and assumptions, making the model less robust than the one for through-the-thickness crack. Chapter 8 summarises the conclusions of the investigations. It can be concluded that with the proposed models, the mechanisms related to the residual strength failure sequence are fully described and characterized. The model for through-the-thickness crack is robust and validated, and can be extended to other material and geometrical configurations. The model for fatigue crack is not robust enough, but further improvements are possible.Mechanics, Aerospace Structures & MaterialsAerospace Engineerin
Design, Analysis, Fabrication and Testing of Grid-Stiffened for Fuselage Applications
No full textThe application of composites in aircraft structures has grown significantly in the past decade, and an increase in this growth in application has recently been achieved in fuselage designs for large civil aircraft such as the Boeing B787 and the Airbus A350. In conventional design philosophies, weight savings is the primary objective on a part-by-part basis, but, with composite constructions, the design philosophies have had to be altered and a direct replacement of a metal with a composite material may not be the most efficient design option. In the future, composite constructions for aircraft fuselage can be made frameless, or at least a reduced number of frames, while making the structure intrinsically damage tolerant owing to its design. One of the front runners in cylindrical stiffened shells for applications that undergo pressurization and therefore, in-plane bi-axial stresses, is the grid-stiffened or lattice structure. A successful application of such structures is the rocket insterstages used in Proton and Delta rockets. A primary advantage of the grid-stiffened construction is the possibility of a long uninterrupted fuselage, which leads to a simple, low-part number, and low-cost design; but the increased weight savings and reduced complexity come with a price: unknown failure modes and the difficulty of joining members to members in an airframe. These challenges are dealt with here.Aerospace Structures and Computational MechanicsAerospace Engineerin
Fatigue and Damage Tolerance of Friction Stir Welded Joints for Aerospace Applications
No full textFriction stir welding is a young welding process with high potential to replace riveted joints in aerospace structures like the fuselage. Friction stir welding is a robust process and capable of welding high strength aluminum alloys. Therefore it can lead to both costs and weight savings. To substitute riveted joints by friction stir welded joints, it must be shown that the damage tolerance of the welded structure is similar as the structure with riveted joints. As the aerospace industry has over 50 years of experience with riveted joints, this is quit a challenge. The damage tolerance behavior of riveted joints is fully understood and the predictability is high. To get friction stir welded structures at the same level of readiness, the fatigue behavior of friction stir welded joints must be fully understood. The main focus of this research was to investigate the fatigue behavior, i.e. fatigue initiation and fatigue crack growth, of friction stir welded aerospace aluminum in order to understand which parameters play a role. This knowledge enables to build prediction models which can be used to evaluate the damage tolerance of a welded structure. Consequently, it is possible to come up with design rules for a damage tolerant friction stir welded fuselage structure.Structural IntegrityAerospace Engineerin
Mode I fatigue delamination growth in composite laminates with fibre bridging
No full textAdvanced composite materials have been commonly used in aerospace engineering, because of their good mechanical properties and attractive potential for creating lightweight structures. Susceptibility to delamination is one of the most important issues in the applications of these materials. This disadvantage can prohibit the application of composite materials in primary aerospace structures and limit their lightweight potential. Therefore, characterizing fatigue delamination growth behavior in composite laminates is important for the applications of these materials in aerospace, as it provides the necessary information for the damage tolerance design philosophy. Fibre bridging is an important shielding mechanism during delamination growth in composite laminates. It can increase the fracture toughness by restraining the crack opening and inhibit delamination growth. However, there is no reliable method to take into account of its contribution to fatigue delamination growth. Thus, investigation of this phenomenon and development of a prediction method is required. The objective of this investigation is to study mode I fatigue delamination growth with fibre bridging in composite laminates and provide physics-based interpretations of fatigue delamination growth. Two approaches are applied to interpret the fatigue delamination growth behavior according to different perspectives. In the engineering perspective, the Paris relation is applied. However, in the physics-based perspective, energy principles are used. The bridging effect on fatigue delamination growth is first investigated by a series of fatigue tests. The Paris relation is used to interpret the fatigue data. It demonstrates that fibre bridging can significantly decrease the fatigue crack growth rate, making it invalid to use a single fatigue resistance curve to determine fatigue delamination with bridging. A new method, still based on the Paris relation, is developed to predict fatigue delamination growth with fibre bridging, by correlating the curve fitting parameters with the amount of bridging fibres. Fatigue delamination growth is physically explained according to the energy conservation law. It is found that bridging fibres periodically store and release strain energy upon loading and unloading cycles. However, they have no contribution to the real strain energy release, unless the bridging fibre pullout or fails. This can lead to the invalidity of using the strain energy release rate (SERR) determined by the fixed grip assumption for quasi-static crack growth to interpret fatigue crack growth. Therefore, the SERR commonly used is not a suitable similitude parameter to determine fatigue delamination growth. Energy principles are subsequently used to interpret the stress ratio effect in fatigue delamination growth. A concept of fatigue fracture toughness is proposed to describe the steady fatigue delamination growth, in which little or no bridging fibre pullout or failure occurs. Therefore, all energy dissipation in the steady delamination growth is concentrated on the new crack generation. The fatigue fracture toughness is observed to be interface configuration independent but significantly stress ratio dependent. The mechanisms related to this are interpreted by fractographic observation. Using the concept of fatigue fracture toughness, the stress ratio effect on fatigue delamination growth can be explained with a clear physical background.Aerospace Structures and MaterialsAerospace Engineerin
Fatigue Crack and Delamination Growth in Fibre Metal Laminates under Variable Amplitude Loading
No full textThis thesis presents the investigation into the fatigue propagation and delamination growth of Fibre Metal Laminates under variable amplitude loading. As explained in the first chapter, the motivation of the research is twofold: first, to obtain a clear understanding and detailed characterization of the failure mechanisms in GLARE under variable amplitude loading (selective and flight load spectra). Second is to obtain an accurate prediction model for fatigue crack propagation in GLARE accounting for fibre bridging, delamination and influence of plasticity. The major concept in this thesis is that the stress intensity at the crack tip in the metal layers of a Fibre Metal Laminates (FMLs) is the factor determining the extension of that crack under cyclic loading. This implies that the stress intensity factor can be described with Linear Elastic Fracture Mechanics, including the contribution of the fibre layers and the with the crack growth associated delamination behaviour. The investigation presented in this thesis, covers the theoretical analysis of the crack growth phenomena and experiments to support and validate the developed prediction model. This investigation has been restricted to through-the-thickness cracks with the same crack length in all metal layers. In the second chapter, the various GLARE grades and lay-ups are defined together with a description of the manufacturing process, quality assurance procedures and fatigue crack growth phenomena. The aspects introduced are the fatigue crack growth in the aluminium layers, controlled by the stress intensity factor at the crack tip, and the delamination of the aluminium and prepreg layers, which occurs in the wake of the propagating crack. The crack opening is constrained by the bridging fibre layers, while the stress redistribution to these fibre layers determines the delamination growth. In addition, this chapter discusses the effects of variable amplitude loading in metals, together with the models developed so far to predict fatigue crack growth. The influence of variable amplitude loading on the delamination growth (i.e. a major phenomenon contributing to the slow fatigue crack growth in FMLs) is discussed in chapter 3. An extensive test programme is presented in this chapter utilizing double and multiple block loads, and flight load spectra to validate the hypothesis that delamination in FMLs is independent of retardation effects under variable amplitude loading. Furthermore, the use of scanning electronic microscopy is presented to study the delamination growth striations on the disbonded fracture surfaces. Chapter 4 presents the effect of variable amplitude loading on delamination shapes. These delamination shapes influence the bridging stresses and the crack tip stress intensity factor. The change in the delamination shape due to variable amplitude loading is understood, and the hypothesis on the reason of this change in delamination shapes has been validated. The use of Digital Image Correlation (DIC) is presented to observe delamination shapes and crack tip plastic zones in-situ fatigue testing. Although a change in delamination shape can be observed, its effect on the bridging stress profile is observed to be negligible. Crack tip plasticity and shear-lip formation both were investigated, as presented in chapter 5. Difference in shear-lip morphology was observed in monolithic metals, adhesively bonded metal laminates and FMLs. Monolithic metals revealed typical shear-lip profiles with the early tensile mode followed by a transition mode and finally the transverse shear mode. For metal laminates (without fibres), the mode transition happened quite late during fatigue crack growth, but the transition length was smaller than those of monolithic metals. In FMLs the shear lip profiles were rather difficult to see with the naked eye or optical microscope. However, the straight appearance of the crack when observed from the specimen side, implied a flat cracked fracture surface without shear lips. Plastic zone sizes are compared in the same chapter. In case of monolithic metals and metal laminates, a big difference is observed between the predicted plastic zone sizes (using Irwin’s relation) and the sizes measured with DIC. However, comparing these sizes for FMLs revealed only a small difference. In FMLs, the plastic zone sizes are observed to be independent of the crack length which corresponds to the constant crack tip stress intensity factors. Chapter 6 presents the development and validation of three different types of prediction models. These models include a linear damage accumulation (non-retardation) model, a yield zone model and a crack closure model. The predictions using these models correlated with experimentally observed crack growth behaviour. Chapter 7 outlines the sub-routine added to the prediction model for post-stretched laminates. This routine has been validated with the data from post-stretched GLARE 1 and ARALL from the literature. Chapter 8 summarizes the conclusions of the investigation. It can be concluded that with the proposed prediction model, the mechanism of crack propagation and delamination growth in GLARE is fully described and understood. The prediction model has been validated with experimental crack growth data, and is considered accurate. In its implemented form, the model has the potential to be extended to other material-, geometrical- and test parameters.Aerospace Structures and MaterialsAerospace Engineerin
Liquid crystalline thermosetting polymers as protective coatings for aerospace
No full textEnvironmental regulations are driving the development of new aerospace coating systems, mainly to eliminate chromates and reduce volatile organic compound (VOC) emissions. Among the various potential options for new coating materials, liquid crystalline polymers (LCPs) are attractive due to their unique combination of mechanical properties and chemical resistance. Their use, however, has been limited mainly due to poor adhesion properties. Thermotropic liquid crystalline thermosets displayed the properties of traditional LCPs, while having the potential to overcome their disadvantages. The present research has been set to investigate the real potential of phenylethynyl terminated liquid crystalline thermosets (LCTs) for coating applications. The coatings were initially manufactured by melt-pressing the LCT resins onto aluminum substrates. This method was selected mainly due to its simplicity and minimal powder requirements. As a first step, the effects of the thermal curing and the molecular weight of the coating resins were investigated. Then, the influence of temperature and molecular orientation were examined. Subsequently, the adhesion and the environmental resistance of the LCT coatings were analyzed. Finally, the applicability of LCTs on aluminum and composite substrates using atmospheric plasma spraying (APS) was explored to address the industrial limitations of melt-pressing (on size, shape, and thermal resistance of the substrate). Being the first approach to the use of these LCTs as coatings, the present work has contributed to the knowledge and understanding of several aspects of the coating. The incorporation of the end-groups was found to promote the adhesion of the coating compared to the thermoplastic Vectra®, while it did not affect the environmental resistance. In addition, the new polymer chemistry allowed the polymers to be ground into a powder suitable for more versatile deposition methods like APS, expanding the range of applications. These LCTs, however, also present several disadvantages. An extra curing step at high temperature is required, during which, the properties are not significantly improved. In addition, the LCTs investigated here become more brittle after curing, which can be a disadvantage for tribological applications. Furthermore, mechanical properties such as the elastic modulus are not significantly higher than those of the thermoplastic LCP; and the formation of aggregate-aggregate interfaces constitute paths for crack propagation. Finally, untreated coating substrate interfaces constitute paths for environmental attack. The main characteristic of these coatings appeared to be their high chemical resistance and low permeability. These coatings are, therefore, applicable for the protection of surfaces exposed to aggressive liquids or flowing gases. Examples include heat exchangers, gearbox housings, undercarriage components, flooring, and hatches. Since these coatings constitute a passive protection, however, the development of a coating system that includes an active protection would be required to extend coating durability.Structural Integrity and CompositesAerospace Engineerin
Distortions and Residual Stresses of GLARE Induced by Manufacturing
No full textGLARE is used in fuselage panels for its improved fatigue and residual strength. In this thesis, predictive models are developed for residual stresses and distortions of GLARE induced by manufacturing. Material behaviour of the epoxy adhesive and the glass-epoxy prepreg is characterised for thermal and curing circumstances. The derived temperature dependent viscoelastic properties can also be used in other applications in which time and/or temperature changes are of concern. The results of this research are important for three main reasons: first, the material thermo-viscoelastic response is derived for the first time and the thermal properties differ from the ones already used in literature for the design of panels made of GLARE. Second, two new models are developed, applicable to any orthotropic material including Fibre Metal Laminates, for cure and thermo-viscoelastic analyses. Third, the predicted residual stresses can improve the design where properties like fatigue and residual strength need to be considered. With the developed models, other effects like cutting, bonding doublers and stringers can be simulated to obtain the results for a completely featured fuselage panel made of GLARE.Aerospace Structures and MaterialsAerospace Engineerin
Fatigue Crack Growth Prediction for generalized fiber metal laminates and hybrid materials
No full textThe excellent durability performance of Glare, a thin fiber metal laminate (FML) material system, is now being proven in service. This has motivated work towards the application of FMLs to thicker structures driven by damage tolerance. In order to fully characterize the crack growth life of such materials, models are necessary that can account for the unique aspects of material systems under consideration, including non-uniformity of composition and stress states, and the resulting complex damage state involved in fatigue crack growth. This thesis presents a generalized analytical model for the prediction of fatigue crack and delamination growth in FMLs of arbitrary lay-up, including differing metal alloys, different thickness layers, and different combinations of reinforcing composite layers. Cracks in each layer, and delaminations in each interface, are allowed to grow separately, with the interactions of the damage throughout the laminate taken into account. The model is structured in a modular and iterative fashion. Modules for determining the load redistribution around damage and the strain energy release rate of delamination have been derived and independently validated through comparison to finite element analyses. A series of tests with thick fiber metal laminates of varied construction was carried out to verify the overall crack growth predictions of the model. While some discrepancies between the results and predictions for the most complex laminates suggest that refinement of the delamination strain energy release rate formulation is needed, many of the test results were accurately predicted, demonstrating the suitability of this model for use in design and analysis of thick FML structures.Structural IntegrityAerospace Engineerin
- …
