This thesis project focuses on the evaluation of the Extended Finite Element Model (XFEM) for modelling skin-stringer separation in thermoplastic composite stiffened panels. The study uses test results from a reference study in the literature on a PEKK-FC carbon composite panel with three stringers joined to the skin using a short-fibre reinforced butt joint. The modelling strategy was developed using Double Cantilever Beam (DCB) specimens. The XFEM-based panel models were then developed using two different damage initiation criteria: Quadratic Stress Criterion (QUADS) and Maximum Principal Stress Criterion (MAXPS). These models were compared with the Virtual Crack Closure Technique (VCCT) model from the reference study. The project also addresses the challenges posed by the crack tip impingement problem in the XFEM-based model with MAXPS and tests different strategies to overcome it. A proof of concept is provided for using the UDMGINI subroutine to address these challenges, setting the stage for future studies and algorithm development.

This dissertation explores the development of a framework for wing aeroelastic tailoring in the preliminary design phase of aircraft, focusing on weight reduction and aerodynamic efficiency through the use of composite materials. The framework utilizes low-fidelity, fast computational tools to handle numerous load cases, modeling the wing structure with FEM shell elements and aerodynamics with the doublet lattice method (DLM). Composite material properties are optimized using lamination parameters, with structural continuity enforced through blending constraints.

To address the limitations of low-fidelity methods, CFD-based corrections are applied to the DLM for accurate aerodynamic load representation, and a surrogate model is introduced for aerodynamic drag estimation. This model approximates wing deflection efficiently, reducing the need for costly high-fidelity simulations.

The framework, implemented in Nastran, is applied to industrial cases focusing on the impact of blending constraints, dynamic load cases, and multi-objective optimizations balancing weight and drag. Blending constraints result in smoother transitions between wing sections, ultimately producing a light final design, even though there is an initial 5% increase in structural weight due to a reduced design space. Dynamic loads, incorporated through the Equivalent Static Load (ESL) method, have minimal impact on the overall design.

Multi-objective optimization reveals trade-offs between structural weight and aerodynamic drag. Weight minimization favors structures with wash-out behavior, leading to higher drag due to increased angle of attack. Conversely, drag minimization results in stiffer wings with less wash-out. The framework's results align closely with full-order CFD models, validating its effectiveness.@en

Offshore installation vessels are growing in size in order that they are capable of handling the next-generation 15 and 20MW wind turbines. At the same time, it is important to gain a better understanding of their structural behaviour using the finite element (FE) method. One approach that is often used is the full vessel FE analysis, which studies the vessel in its entirety. However, the drawbacks are that a coarse mesh must be adopted to have an acceptable simulation time and it takes months to create a 3D model of a vessel. Another approach is the partial vessel FE analysis, which considers only a section of the vessel. This research aims to develop a methodology for the partial vessel FE analysis of an offshore jack-up installation vessel.

The studied section of the vessel is between the four legs, which is where the highest bending moment is expected and the wind turbine towers are located on deck. The methodology was developed to analyse the vessel in still water and waves separately. The boundary conditions and hull girder load adjustments of the partial vessel model were described, which cause the vessel section to reflect the behaviour of the full vessel. The procedure to apply the loads on the model was also given. The finite element analyses were performed and the stress results compared to a simple analytical solution of the vessel based on Euler-Bernoulli beam theory.

The methodology of the partial vessel FE analysis was successfully applied to the offshore jack-up installation vessel. For the vessel in still water, the stress results were in agreement with the analytical solution. The results were accurate in the middle region for approximately 50% of the partial vessel model and thus it is advised to have a larger model for future use. The effect of torsion could not be fully justified as it is not captured in the analytical solution, but resulted in a significant improvement. For the vessel in waves, the stress results also showed a good agreement with the analytical solution. The accuracy depended on the equivalent design wave approach, which assumed a regular travelling wave. Lastly, it was proven that the presence of payload only leads to local changes in the stress field, but not in the global behaviour of the vessel.

The issue of space debris in Low Earth Orbit is a growing concern that requires comprehensive solutions. This includes preventing further pollution, removing existing debris, and designing resilient spacecraft that can withstand impacts. This thesis focuses on the improvement of spacecraft shielding structures to provide effective protection against hypervelocity impacts. The Smoothed-Particle Hydrodynamics method was used for the simulation of hypervelocity impacts using LS-DYNA. The method is preferred for the simulation of impacts at high velocity and was used to reproduce two studies from the past. The simulation was then performed on a double plate system, with results compared to data from experiments provided by Airbus Defence and Space GmbH, under co-supervison from ESA. The simulation accurately predicted the hole diameter in the 1st plate with less than 2% error, while the damage zone in the second plate showed considerable variance. The simulations consistently overpredict the damage zone, leading to conservative results.

The rudder is one of the few aircraft primary composite structures that experience relatively low external aerodynamic loading during regular flight conditions. This renders the rudder shell the perfect candidate to study the possibility of introducing sustainable materials with lower mechanical properties in primary structures. However, high internal loads are introduced in the shells during actuator jamming events, which occur with low frequency. Locally redesigning the rudder spar as a torsion box to take up the jamming loads and alleviate the shell loading could enable the safe usage of green materials in the rudder shells. The aim of this thesis is to study the design feasibility of three rudder torsion-box spar concepts to separate the aerodynamic and jamming loads in the rudder and provide a preliminary optimised structural design based on a parametric study of the torsion-box spar design variables and static FEM simulations.

This thesis aims to study the design and optimization of composite wing components for aircraft, focusing on the pivotal goal of reducing weight while preserving and enhancing structural performance. This is one of the goals of the COST Action joined in the span of the thesis: CA18203 - Optimising Design for Inspection (ODIN). Within this COST Action, a comparison between different Finite Element modeling predictions of the response of this structure subjected to 4-point loading has been presented
by Bisagni et al. [1] before this thesis. The thesis, which corresponds to the optimization of this same structure, is the continuity of this work. The foundation of this thesis is built upon a literature study that investigates the optimization of composite wing structures. It explores diverse optimization formulations, including design variables, objectives, and constraints. Different methodologies for optimizing wing structures, like multi-objective and probabilistic methods, are studied. Various algorithms used to tackle these challenges are presented, shedding light on their advantages and drawbacks. Once a solid knowledge of wing structure optimization is built, preliminary analyses (study of coupon geometries, failure criteria, and the buckling behavior of various materials and geometries) form the basis for the upcoming optimization strategy. These preliminary investigations refine our understanding and guide the selection of optimal constraints, such as failure criteria. Based on these preliminary analyses and the literature study, the optimization strategy has been established. A new configuration with cut-outs is developed to explore new innovative designs. To solve the optimization problem, Genetic Algorithms, inspired by real-life behaviors and processes, emerge as the optimization algorithm of choice. These algorithms contribute to the development of design solutions that are better than traditional designs. The objective is to achieve weight reduction in the structure while considering its structural performance. Consequently, the fitness function, employed to assess candidate designs within the Genetic Algorithm, is based on three primary factors: weight, stiffness, and buckling behavior. Weight carries the most significant part of the fitness function, accounting for approximately 80% of the total value, whereas the remaining two factors contribute approximately 10% each. To evaluate these aspects, weight is determined through basic calculations, while stiffness and buckling behavior are assessed by simulating a 4-point bending test using Abaqus.
The implementation of the Genetic Algorithm, adapted to our special case, allows us to showcase its effectiveness in achieving optimal designs. These optimal designs achieve remarkable weight reductions while maintaining great structural performance. Following 20 generations, each comprising 10 potential candidates, one of the two termination criteria is triggered: there is no improvement in the best solution for eight consecutive generations. The optimized solution obtained is 16.89 kg lighter compared to the baseline configuration studied by Bisagni et al. [1]. This represents a significant 25% reduction in structure weight. Concerning the structural performance, only a reduction of 1.5% and 0.25% is observed for the stiffness and buckling performance, respectively, which is acceptable regarding the weight reduction. Nonetheless, achieving such a reduction in wing weight may not be realistic as it was specifically tailored for the test case of this thesis, other loading scenarios need to be studied to assess its applicability.
These outcomes underscore the potential of optimization techniques in aerospace engineering, paving the way for the development of lightweight and high-performing composite wing components.

The state-of-the-art design approach of composite structures in Formula One (F1) rely on expensive full-scale crash tests and simple models that cannot describe the underlying physical phenomena during crushing. Although the models can describe the crush forces and resulting accelerations, their predictive capabilities are poor, and several iterations are usually required to arrive at a satisfactory design, causing the teams money and resources. Reductions in costs could be possible by reducing the number of full-scale experiments and increasing the amount of analysis.

This thesis aims to evaluate the capabilities of a commercially available material model at two distinct modelling scales by conducting physical and virtual experiments on two different specimen geometries; tubular coupons and a new open-section channel specimen. Firstly, the new open-section specimen with a geometry representing the typical F1 crash structures was designed, manufactured and tested. The newly proposed specimen showed controlled failure modes, and displayed typical characteristics of composite structures under crushing loading. Secondly, the two specimen geometries were numerically analyzed on the macro-scale (laminate-level) and meso-scale (ply-level) using a commercially available material model implemented in the Abaqus finite element software. Mesh-dependency of the macro-scale approach was discussed and numerical aspects required to perform crush simulations on the meso-scale were identified and discussed. Lastly, calibration of the numerical models was performed to understand the effects of model inputs on the output responses, to evaluate the validity of the input parameters, and to assess two different calibration approaches; the first method is based on macro-scale modeling, while the second method is based on surrogate-based inverse parameter identification at the meso-scale.

The numerical models could not reproduce the experimental crush responses of either geometry when using experimentally characterized material data. Following the second calibration approach, the qualitative and quantitative crush responses of the tubular specimen could be accurately described, but the channel specimen showed no improvements in their response. The macro-scale modelling approach with the current material model was judged to be impractical for large-scale analysis, and the meso-scale approach showed need for further developments.

The safety of a commercial aircraft is a factor to consider from the early stages of its design to guarantee that this one's structure can protect the occupants inside in a low-speed crash. In an aircraft's primary structure such as a fuselage, where the occupants and cargo should be protected, the design regulations imposed by the EASA have been applied for decades to conventional metallic designs. These designs had aluminum as the main constituent for being a lightweight material. Nevertheless, this project improves a component's design from a fully thermoplastic fuselage. While two commercial aircraft (A350, and B787) have already implemented some composite parts in their design to aim for sustainable and lightweight components, thermosets were the composites used in primary structures mainly. Composite materials do not present the plasticity metals do, which means that the regulations from EASA can no longer apply in the A350 and B787 hybrid designs. With the introduction of composite materials in these aircraft, organizations like the FAA and EASA had to write some new guidelines for these aircraft structures. These enable the protection of the passengers and cargo from the loads exerted on the aircraft during a survivable crash. Therefore, the new regulations are the ones taken into account for this thesis project, since these can be applied to similar composite structural designs, such as the fully thermoplastic fuselage section, whose numerical model is inherited from the Clean Sky STUNNING project.

What is wanted for the fuselage section's crashworthiness is for its structure to absorb as much energy as possible from the crash. All while minimizing the peaks of force that this one experiences. That way, the loads that arrive at the cabin of the aircraft are reduced and the structure becomes safer for a low-speed crash of under 30 ft/s (about 10 m/s), meaning that the passengers inside have a higher chance of survival and it is less probable they suffer from severe injuries. That is the aim of this project, focusing only on the structure's behavior caused by the changes in the fuselage struts.

The numerical analyses performed in LS-DYNA increase single components’ structural complexity until optimizing the layup of a tube to improve its crashworthiness behavior. It is after this optimization that the component is introduced as a strut in the STUNNING fuselage section. To further improve its global crushing in low-speed crushing conditions, the struts (or energy absorbers) keep the square geometry cross-section from the tube exercise, as well as the optimized layup. However, these are changed in size to improve the fuselage’s crushing behavior with little to no increase in mass. This study considers a fuselage section that neglects the passengers’ mass, their cargo, and the upper part of the fuselage airframe in the section's model. Considering only the cargo mass for six passengers on the model causes a change in the fuselage’s final design crashworthiness, which proves that a more representative crushing conditions setup should be considered in future studies to better predict the structure’s crushing behavior numerically.

A novel twist morphing concept is explored, that exploits the buckling instabilities of slender spar webs integrated into a wing structure to control the twisting response to external loads. The main novelty of the concept is in controlling the effective shear stiffness of the post-buckled slender spar webs through external variable constraints acting on the out-of-plane buckling deformations. A methodology for the design and analysis of these novel morphing structures is proposed and implemented in the design of a wing box structure of promising twist morphing capabilities. The overall design process is structured into a multilevel process of increased complexity. In the first level, the morphing structure is simplified to a wing box with slender spar webs. With the objective of maximizing the morphing twists that can be achieved under the action of an external quasi-static torque, the wing box design space is explored in terms of its cross-sectional dimensions and the material assigned to the slender spar webs. In the second level, the morphing structure is expanded to include both the wing box and the external devices required to implement the adaptive constraints acting on the slender spar webs' out-of-plane buckling deformations. At this level, the objective is to design adaptive constraining devices that maximize the twist morphing capabilities, for which the influence of the constraining devices over the twisting response and their effectiveness in restraining the slender spar webs' buckling deformations become the main concerns. After an extensive design process, a design solution for the adaptive constraining devices is proposed, for which thorough analyses on the twist morphing capabilities are performed. In addition, the sensitivity of the twist morphing capabilities to the slender spar webs' geometrical imperfections is investigated.

The goal of this thesis is to build an artificial neural network(ANN) surrogate model, that predicts the crashworthiness performance of a structure. The structure used in this research is a thermoplastic fibre-reinforced composite aircraft subfloor section which is part of the SmarT multifUNctioNal and INteGrated thermoplastic fuselage (STUNNING) fuselage demonstrator section. The structure will be analysed numerically, by making use of the finite element software LS-Dyna by Ansys. Within this research, a surrogate model is defined as a model that has been fed data from a computationally expensive model, yet is computationally considerably more inexpensive to run to obtain similar results as the expensive model. The main methodology to do so in this research is to build a two-dimensional design space, which is sampled using optimal latin hypercube sampling. By using an existing model for the STUNNING fuselage subfloor section and by automatic generation of geometry changes of the structure through a Python script, input files can be created according to the sampled design space. The simulations are then run on a high performance cluster. Once the simulations are run the results can be fed to an artificial neural network that uses this data to predict crashworthiness performance of the structure within the design space.

To build towards the goal of the thesis, a small comparative study has been performed on material models MAT054, MAT058, and MAT 261 in LS-Dyna. These material models are based on different failure criteria, and thus have different performance in coupon simulations. Using these simulations as a basis, a simple design space was created, sampled, and modelled so that a first ANN surrogate model could be built that could predict the failure displacement and the ultimate load of a composite coupon in tension. It was found that for a small dataset (<10 samples), the neural network could easily fit to the data, and could predict accurately on samples within the design space that were withheld from the network during training.

Then, the STUNNING subfloor section model was introduced and a number of models that section this structure are investigated. The main purpose of this investigation is to find out if the model can be reduced in size without changing the resulting failure mode. When a satisfactory reduced model was found, the research carried on by doing a sensitivity analysis on a proposed two-dimensional design space, and similarly a sensitivity analysis on the parallelization settings. From this study, a second failure mode of the model was discovered and the effect of the parallelization settings was quantified. From here the limits of the design space could be formed and sampled. Finally, different variations of ANN surrogate models were defined, trained, and tested. In the end, a root mean square discrepancy of around 10\% was achieved during testing of the network. This means that the methodology could prove useful in design applications, depending on the design stage and needed accuracy.

In recent decades, fibre-reinforced polymer composite structures have seen increased aerospace applications. A typical composite structural element widely used in primary load-bearing aircraft components is the T-joint, which is formed when two orthogonal laminates are joined in a T-shaped arrangement to enable out-of-plane load transfer. Under tensile loading, the structure tends to fail by the separation of the horizontal and vertical laminates at the joining interface. The incorporation of an interleaf, i.e. a thin layer of material with good interfacial bonding properties, may prevent or delay the failure event. A potential interleaf material is the carbon nanotube (CNT). In the literature, its addition as a reinforcing phase in interleaves has been shown to improve fracture toughness in standard coupon tests, but the technique has yet to be tested on a structural element level.

This thesis research investigated the effect of CNT interleaves on the mechanical behaviour of composite T-joints. Manufacturing processes were developed to produce test samples consistent in quality. T-joint samples with CNT-modified interleaves were manufactured alongside neat samples with zero CNT content. The joints were formed using custom-built tools and co-cured in an autoclave. Tensile tests were carried out and the T-joints were loaded until failure. Load-displacement data were acquired and further supported by results from a Digital Image Correlation (DIC) system.

Data analysis suggested that CNT interleaves modified the mechanical behaviour of T-joints by delaying the failure of the structure. While their stiffness and ultimate strength were largely unaltered, the CNT-modified T-joints were able to undergo higher displacements and absorb higher strain energy before failure compared to the neat samples. The delay of failure was enabled by extensive crack development within CNT interleaves preceding the final rupture. In comparison, the neat samples failed abruptly after very limited damage evolution. Observation of the crack surfaces revealed evidence of tortuous crack paths through CNT interleaves. The finding suggests that CNT interleaves are capable of improving the failure resistance of the T-joint structure in a targeted fashion.

The collapse of composite stiffened panels in compression is complex and involves the interactions of many different failure modes. However, when the panel is allowed to enter the post-buckling regime, skin-stringer separation due to the interaction of the post-buckling deformations with the skin-stringer interface is often the critical failure mode. The accurate prediction of skin-stringer separation in these types of panels is therefore crucial for their design. Due to the high computational costs associated with modelling damage in a large multistringer panel and the manufacturing and testing costs associated with testing such a panel, single stringer specimens that accurately represent the behaviour of critical regions in these panels have been proposed in literature. In this thesis, transversely loaded single stringer specimens are designed and verified and are used to predict skin-stringer separation in the critical regions of a specific multistringer panel. A specific composite stiffened panel is considered that contains a skin of a tape material and four co-cured stringers of a fabric material. At the intersections of the skin and stringers, resin-rich noodle regions are created. A buckling analysis is first performed on the panel to obtain its post-buckling deformations. Based on these, two types of regions where skin-stringer separation may occur are identified; a mode I dominated skin-stringer separation at the minima of the buckling waves and a combined mode II and mode III skin-stringer separation at the inflection points of the buckling waves. A seven point bending specimen is designed based on the panel deformations near the former critical region, while a four point twisting specimen is designed based on the deformations near the latter region. Finally, detailed damage models are set up for the single stringer specimens to predict skin-stringer separation and the models are verified with the multistringer panel. The analyses showed that a mode I dominated skin-stringer separation is critical for this panel. The detailed analysis of the four point twisting specimen showed that no damage occurred at the inflection point. To predict the skin-stringer separation at the minimum, a verified seven point bending specimen has successfully been obtained. To do so, it was important to accurately simulate the deformation of the skin at the stringer flange edge, since this proved to be driving for the initiation and propagation of skin-stringer separation. Next to that, a trilinear cohesive law was used to capture the R-curve effect in the tape/fabric interfaces. Finally, a first assessment of the noodle region behaviour showed that damage in this region will probably happen at the location of maximum bending. Recommendations for future research have been provided.

Launch vehicle structures are commonly composed of cylindrical and conical shells, which are inherently sensitive to buckling. The axial compression experienced during launch can consequently be a sizing load case, so it is important to understand the axial buckling behavior of these shells. Experimental testing is an essential part of studying this phenomenon because manufacturing imperfections can cause large discrepancies between theory and reality. In addition, conical shells have been researched less frequently than cylindrical shells, such that their behavior is less well understood. Experimental testing of launch vehicle structures is difficult and expensive due to their large size, hence it is preferred to test reduced-scale shells, representative of the full-scale ones. In this thesis, a scaling methodology is developed which allows designing representative reduced-scale conical shells for full-scale composite conical shells buckling in axial compression. The conical shells are assumed to have a symmetric, balanced layup with negligible flexural anisotropy. The scaling methodology is developed using the nondimensional governing equations, obtained through Nemeth's procedure, which allows to directly use the coefficients of the equations as scaling parameters. It also provides a framework to not only compare the buckling load of the shells of different sizes, but also the displacement upon buckling, the deformation shape, and the radial displacement. The methodology is set up such that the reduced-scale design parameters are determined sequentially. The buckling behavior of the two shells is compared using a semi-analytical approach, linear eigenvalue, and implicit dynamic finite element analyses. The eigenmode imperfection sensitivity is also evaluated. The methodology is successfully applied to isotropic, cross-ply, quasi-isotropic, and sandwich conical shells. The prediction accuracy is mainly affected by not being able to simultaneously satisfy all scaling parameters, by non-negligible flexural anisotropy and transverse shear, and by differences in imperfection sensitivity between the full-scale and reduced-scale shells. In any case, accurate results are obtained for the considered shells. The radial displacement is most difficult to predict, which is attributed to the membrane prebuckling assumption and neglecting the presence of imperfections. Finally, it is observed that larger eigenmode imperfections affect the accuracy, but they do not cause the methodology to fail. For future work, it is recommended to validate the methodology through experimental testing.

This thesis aims to provide more insight into the crashworthiness behaviour of a composite aircraft fuselage. This is achieved by studying both analytically and numerically the crushing behaviour of composite energy absorbers. The analytical model, which is based on energy dissipation rates, generalises previously derived analytical models to study a wider variety of structural components. It is found that the analytical model can give an estimation of the mean crushing load of square, C-shaped, and corrugated specimens. Next to the analytical study, numerical models are created to gain insight into the behaviour of the studied shapes when subjected to crash loading. Another part of the numerical study investigates the effectiveness of different material models to capture the complex failure mechanism of the composite crash absorber. In LS-Dyna, the MAT054, MAT058, and MAR262 material models are used, which are all able to reproduce the results obtained from reference tests. The ability of the material models to recreate the crushing phenomenon is thought to originate from the opportunity LS-Dyna offers to degrade material properties in the crush frond, which incorporates the formation of cracks and delaminations. In the final stage of this research, a case study is performed, where different absorbers are introduced into a simplified digital twin of the Next Generation Multifunctional Fuselage Demonstration as developed by the STUNNING project. Here, it is found that the fuselage's energy absorption can be increased by introducing energy absorbers. All the simulated fuselage sections show similar behaviour in the first crash stages, characterised by the flattening of the lower section of the frame. Subsequently, the behaviour of the absorber and the surrounding structure highly depends on the absorber's integration. Finally, it is found that by combining the numerical results of a baseline fuselage section and the analytical model for the energy absorbers, one can estimate the energy absorption of the augmented structure with a discrepancy of less than 20$\%$. This proves that the suggested analytical-numerical method can aid engineers during the preliminary design for crashworthiness of a thermoplastic composite aircraft fuselage.

Predicting the critical buckling load of cylindrical shells with circular cutouts subjected to uniform axial compression is an important part of the structural design in the aerospace industry as buckling significantly reduces the load-carrying capability of the structure. A cutout constitutes a major disruption in the shell geometry, and therefore it should be expected that it has a significant effect on the sustainable buckling load. An analytical solution for estimating the buckling load of isotropic and quasi-isotropic composite cylindrical shells with circular cutouts is developed to assess changes made to the geometry and the material during the preliminary design phase quickly. The Ritz method is employed to minimize the total potential energy of an ideal shell that contains a central opening in order to predict a linear buckling load. Finite element simulations are conducted to verify the accuracy of the analytical solution. In addition, they are used to investigate the evolution of buckling modes, the effects of initial geometric imperfections, as well as the shell failure mode. The nondimensional curvature parameter α can be used to categorize the buckling behavior of cylindrical shells and is a function of the cutout radius, the shell radius, and the shell thickness. A small cutout has virtually no influence on the buckling load compared to a pristine shell and the displacement pattern at buckling is global. The buckling load decreases rapidly for moderately large cutouts where the stability loss is the result of a local buckling mode that immediately leads to global buckling. Cylindrical shells with large cutouts are again relatively insensitive to an increase of the cutout size, but the buckling load is greatly reduced relative to a shell without a cutout. Large openings also feature a stable local buckling mode where substantial lateral prebuckling displacements emerge before the structure buckles globally. While the analytical procedure theoretically should not capture the onset of global buckling independent of local buckling, it follows numerical trends for cutouts of moderate and large size regardless. Therefore, it may be used during preliminary design to estimate the impact of changes made to the shell geometry and material. Local buckling is caused by high compressive stresses next to the cutout and, in some cases, large lateral prebuckling displacements. The detrimental effect of the stress field may be partially relieved in composite cylindrical shells by reducing the amount of axial bending stresses that occur. Hence, the chosen stacking sequence can have a significant influence on the buckling load of shells with moderate and large openings.

As collisions between aircraft and birds happen many thousands of times every year, all aspects of such impact events need to be taken into account in the process of designing an aircraft. One aspect that poses a risk is caused by the vibrations induced during bird impact. These vibrations are propagated through the aircraft structure by stress waves. Severe vibrations can negatively affect aircraft components and electrical systems located near the point of impact. While a vast amount of knowledge and research on the modelling of direct bird impact damage is already available, a gap in understanding exists regarding such vibrations. The purpose of this thesis is to help fill this research gap. Therefore, this thesis tries to answer whether the shock effect caused by the stress waves generated during a bird impact on an aluminium alloy plate can be accurately predicted using state-of-the-art numerical modelling techniques. Two aspects are investigated. Firstly, it is evaluated whether the stress waves that propagate the impact loads through the structure can be identified in the model. Secondly, it is analysed how well the severity of the induced vibrations can be modelled. This is done by utilising five bird impact tests, which were performed by Airbus. In each of these tests, a bird impacts an aluminium alloy test plate, which represents a simplified aircraft structure. A dummy mass, attached to these test structures, represents an electrical system influenced by the vibratory response of the plate. The displacement of the plate is recorded using digital image correlation (DIC). Accelerometers at several locations of the test structure are used to record the vibrations. The author created a dynamic explicit finite element model of the test structures to analyse how well current modelling techniques can determine these stress waves and vibrations. The structure in this model employs mostly conventional shell elements. For the bird, a smooth particle hydrodynamic (SPH) model is used. Several types of stress waves are analysed. The main stress wave types evaluated are the in-plane longitudinal and the out-of-plane Rayleigh wave. An attempt is made to identify these waves in both the numerical model and the physical test. It is found that the DIC data of the impact test shows too much disturbance to evaluate these waves. In the numerical model, both an in-plane and an out-of-plane stress wave is seen. Their recorded velocities are seen to approach their theoretical value as the mesh is refined. The propagation velocity of the out-of-plane stress wave is found to depend on the structure of the mesh. When propagating diagonally to the mesh, a 17% lower wave propagation velocity is found compared to a wave travelling tangentially to the mesh. No research was available in which the vibrations of an aircraft structure due to bird impact are analysed. Therefore, the shock response spectrum (SRS) technique, as applied in the space industry to evaluate the severity and damaging potential of vibrations, is utilised. This is the first known research in which an SRS is applied to a bird impact problem. In this research, the SRS is used to compare vibrations from the tests and the numerical model. Such an SRS is generated from the acceleration-time signal of a vibration. It is found that the trend of the SRSs obtained from the numerical model approximates the SRSs from the bird impact test closely. However, due to the fickle pattern of the SRSs, a maximum discrepancy of 5 to 10 dB between model and test is common.

So far thermoset based composites have been used in primary aircraft structures, owing to their good performance and vast research on the subject. Recently thermoplastic based composites also gained an increased attention and there is a push towards adopting these for primary aircraft structures as well. This interest for thermoplastic composites in aeronautical structures is driven by the material's advantages over thermoset parts in joining methods, damage tolerance, sustaining, shelf live, recyclability, or chemical stability. The butt-joint feature is one of the novel joining techniques made possible by using thermoplastic composites, joining technique which can used to assemble stiffened panels by co-consolidating different, typically flat, laminates. With the potential use of thermoplastic structures in primary aircraft structures being relatively recently considered, the available literature on the damage behaviour of thermoplastic stiffened panels under compression is still scarce, to say the least, especially when using this novel butt-joint feature.

In this project, the building-block approach is used to develop a robust modeling technique able to model skin-stringer separation in butt-joined thermoplastic panels under compression. Here the modeling strategy is developed in a step-by-step fashion, going from a material characterisation coupon scale, up to a sub-component scale. The separation was first modeled in a DCB coupon used to describe the Mode I material fracture toughness, then in a single-stringer specimen with skin-stringer de-bond and finally in a multi-stringer panel with skin-stringer de-bond. The base of the damage modeling strategy was defined by studying the load-displacement response of a DCB specimen using multiple models. First, an experimental load-displacement response of a DCB specimen was captured by an analytical approach, with which the accuracy of the numerical models was compared. The numerical approaches tried were based on cohesive zone models (contact and element based) and on the Virtual Crack Closure Technique, with the latter proving to be better suited for this particular task. The VCCT separation damage model was preferred here since it allowed the use of coarser meshes than the cohesive zone models, with a superior overall accuracy as well. As the material of the studied DCB specimen was identical with the one identical of the single-stringer specimen and multi-stringer panel aforementioned, the VCCT model used to model separation was also kept identical when these larger structures were studied. The accuracy of the FE model developed for the single-stringer specimen was also assessed by comparing its load-displacement curve and failure load with the existent experimental data, improving the modeling technique for its use in the multi-stringer panel. As an FE model which correlated well with the experimental data already existed within Fokker, using a cohesive surface based damage model, the shortcomings of this damage model was shown by applying it to the lower scale DCB specimen. Implementing this damage model for the DCB specimen led to a very poor load-displacement correlation, owing to the altered fracture toughnesses values with respect to the measured one. With the shortcomings of the old separation damage model shown, the VCCT one was implemented in this already existent model, which gave a poor correlation with the experimental data. Since the VCCT damage model used needed to be kept unmodified, the butt-joints of the single stringer were refined and the elements used for these areas were updated in a new FE model, these changes leading to a near excellent experiment-simulation correlation in terms of load-displacement curves and failure loads. The modeling technique developed was then used further to study skin-stringer separation in the multi-stringer panel. A sensitivity study on the panel's compressive response with and without the de-bond damage modeled revealed the high sensitivity of its skin buckling pattern. The occurring skin buckling pattern appeared to be highly influenced by the skin's imperfect non-flat shape, asymmetric load introduction, ends boundary conditions, as well as by the de-bond damage itself. This very sensitive skin buckling pattern was attributed to the specific combination of panel's skin bays boundary conditions and skin bays aspect ratio, which likely gave a buckling coefficient governed either by a 3 or 4 longitudinal buckling half-waves. The sensitivity study on the different possibly occurring buckling patterns revealed that the ones promoting de-bond growth from a certain side of the stringer have a strong influence on the panel's failure load, decreasing the strength of the panel by as much as 25% for a 4 half-waves skin buckling pattern, the most likely to occur however having a 3 half-waves one. Similarly, the most significant strength increase was also due to a 4 half-waves skin buckling pattern, the difference between this and the aforementioned one being the stringer side from which the de-bond grew. Furthermore, this sensitivity study was also used to select a best blind prediction based on which the experimental test plan was defined. The test measurement set-up to correlate and validate the developed FE model was also defined based on this selected best blind prediction and it was successfully used during the experimental test to accurately capture the panel's compressive response. The test-model correlation showed that the FE model was able to predict the compressive response of the panel with great accuracy in what concerns its load-compression response, failure load, buckling behaviour, skin out-of-plane deflections and the qualitative aspect of the de-bond growth. These qualitative aspects of the de-bond growth were its skin buckling pattern influenced location, de-bond growth front shape and its growth soon after skin buckling occurred. A initially very good correlation in terms of load-strain response was also achieved, this correlation being highly influenced by two aspects. First source of test-model strain correlation mis-match was considered to be the significant difference in the panel's skin buckling load, while the second one was the model's poor prediction of the panel's stiffness drop associated with de-bond growth. The de-bond growth was overestimated by the FE model prediction, while the panel stiffness loss as a result of the de-bond growth was underestimated. Overall, the presented modeling strategy to model and study skin-stringer separation in thermoplastic butt-joint stiffened panel under compression proved to be robust. The robustness of this method comes mainly from the same modeling strategy used at all the addressed scales and with the same VCCT damage model using the measured material fracture toughnesses. This implies that the presented method could be successfully transferred to similar panels with similar skin-stringer damage, provided that skin-stringer separation is the main failure mode.

Cylindrical shells and conical shells have been analysed and tested separately for the last decade. However, due to new manufacturing techniques and the use of composite material systems more complex structures can be made. By combining a cylindrical shell and a conical shell one structural component can be created: a cylindrical-conical shell. This is a potential design solution for launch vehicle structures, but there is limited research available about it. Not much is known about the buckling behavior or imperfection sensitivity of these structures, especially not for sandwich composite shells. The objective of this research is to understand these two phenomena in cylindrical-conical structures, considering a sandwich composite. The buckling behavior is studied by using finite element analysis and the behavior of the individual shells are compared to the combined structure. Different shell geometries such as height of the cone, angle of the cone, and radius of curvature at the transition are considered. The imperfection sensitivity of the different shells is investigated by adding axisymmetric and eigenmode imperfections using different amplitudes. The results indicate that the buckling behavior of the individual components without initial imperfections can be captured by the linear eigenvalue analysis as well as by a postbuckling analysis using a dynamic implicit solver with non-linear geometry enabled. In case of the cylindrical-conical shell, the linear eigenvalue analysis is not able to capture the buckling shape and results in an overpredicted buckling load. This can be explained by the non-linear behavior at the transition area before buckling that can be observed in the dynamic implicit analysis. This non-linear behavior before buckling is mitigated when a reinforcement or radius of curvature is added to the cylinder to cone transition region. A similar sensitivity to geometrical imperfections is observed for the cylindrical shell and conical shell. The cylindrical-conical shell has a much lower sensitivity, because the non-linear behavior at the transition region is dominating the buckling behavior of the whole structure. The results obtained from the thesis show the importance of the non-linear behavior at the transition for the sandwich composite cylindrical-conical shell. This means that for further research it is important that the non-linear effects are included when analysing such a structure.

Shape morphing has been considered a future direction in the pursuit of maximised aerodynamic and structural efficiency of an aircraft wing by triggering geometric reconfiguration to adapt to specific design requirements or load environments. Researchers have endeavoured to seek for morphing solutions that are simple, predictable, tailorable and with promising shape adjustment. Instead of the traditional structural design against buckling failure, a novel concept now embraces this built-in instability by utilising the nonlinear buckling response of slender elements to produce selective stiffness modification which redistributes the load in the structural system. To enable desired buckling configurations with specific stiffness, three buckling-driven mechanisms are presented in this thesis by restraining the out-of-plane buckling deformation with point, area and maximum displacement constraints. Numerical investigation of the proposed mechanisms is first conducted on a composite plate to analyse the effect on the onset of buckling and postbuckling stiffness modification. The laminate with buckling-oriented mechanisms implemented is then integrated into a thin-walled rectangular composite wing box as a shear web to evaluate its functionality in a structural system. The proposed mechanisms offer effective control of the configurations and augmentation on the attainable range of shear centre relocation and torsional stiffness variation of the wing box. To explore the feasibility in morphing application, the interaction between buckling-induced twist and the aerodynamic load distribution is studied with a simplified aeroelastic wing model. Served as proof of concept, the work of the thesis demonstrates the potential to realise complicated morphing tasks through employing precisely-controlled buckling behaviours in structural components.

The structural properties of a composite laminate can be tailored by allowing the fibres to vary their orientation within the individual plies. This class of composites is termed variable stiffness and they are able to enhance the structural performance over traditional laminates by redistributing the internal load of the structure. Therefore, it is of interest to pinpoint the structural properties that may benefit most from this concept and the extent thereof. To that end, a multi-objective optimisation of a variable stiffness cylinder was performed which identifies the Pareto front for objectives such as buckling, natural frequency, stiffness and strength. Manufacturing considerations were also incorporated by accounting for the allowable tow curvature and the resulting overlaps during the process. The stiffness variability was formed by means of a Lagrange polynomial, the non-linearity of which is determined by a specified amount of control points. An investigation on the suitable degree of non-linearity of the functional representation of the cylinder was thus carried out. In order to complete the optimisation in a practical time frame, machine learning algorithms were utilised to act as surrogate models and perform predictions on the optimisation objectives. Different algorithms were implemented for that task, namely Radial Basis Function, Kriging and Artificial Neural Network models. The Artificial Neural Network surrogate model outperformed the others when the dimensionality of the problem was high, although the opposite was observed for low dimensional problems. Cylinders with and without cut-outs were optimised separately and compared with their optimised constant stiffness counterparts. It is found that the performance gains for the cylinder without cut-outs are not significant when the additional mass due to the overlaps is accounted for. On the other hand, the cut-out cylinder greatly benefits from the stiffness variation for multiple objectives. Additional structural improvements were witnessed for higher degrees of non-linearity of the utilised Lagrange polynomial.