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.

Guidelines dating back 50 years, NASA SP-8007, are employed today in the design of thin-walled launch vehicle structures. Due to advances in materials, structural designs, and manufacturing techniques since the publication of SP-8007, the development of new knockdown factors for contemporary launch vehicle structures is an ongoing subject of research. The work presented herein was performed in collaboration with the NASA Engineering and Safety Center on the Shell Buckling Knockdown Factor Project. A laboratory-scale composite cylindrical shell test article, which had previously been designed according to a novel scaling methodology, was the subject of simulation and testing. Its inner, outer, and boundary surface imperfection signatures were measured and implemented in finite element models for buckling test simulations. These were then used to provide prediction data for an experiment conducted at NASA Langley Research Center. Buckling loads from the two pre-test analyses were within 0.08% and 3.7% of the experimental buckling load. The concurrence of axial shell stiffness, localized strains, and buckling shape evolution was also established between the experiment and simulations. A slight loading imperfection was found during the test; however, it was demonstrated through post-test analyses that this did not affect the buckling load substantially. The test article's 0.91 normalized buckling load was much higher than the 0.59 knockdown factor specified by SP-8007. The correlation between the experimental and simulation results, as well as their contrast with SP-8007's prescription, suggests that directly measured imperfections are capable of playing a role in the development of modern and potentially less conservative knockdown factors for future launch vehicle structures.

The Europa Lander is a mission by NASA to land scientific instruments on the Jovian Moon by 2030. The current design of the landing system envisions fixed landing legs with the lander being lowered down at very low velocities (≈ 0.1 m/s) by use of a Skycrane. This latter system not only adds mass, but also requires complicated moving parts and a high reliance on control systems to work properly. This thesis wants to propose an alternative design to the Europa Lander Landing System that uses sturdier legs that can withstand landing at higher velocities than originally envisioned. This eliminates the need to use a Skycrane system. In order to pack them in a more efficient way while coasting, the landing legs are of the deployable type. In order to track the deployment dynamics of the landing system a 2-dimensional Matlab deployment model based on double pendulum motion with damping is created. Furthermore, in order to prove the landability with a prescribed landing case, a landing model is created using FEA software Abaqus. The landing model is used to make sure that stresses, strains and G forces applied to the rest of the lander don’t exceed prescribed values as dictated by requirements. The implementation of both models is verified by comparing models found in literature with the ones made for the purpose of this thesis. Good correlation is shown with less than 10% of difference between literature results and results for this thesis. The result through the modelling found a specific geometry with parameters that include, but are not limited to: length of leg elements, joint damping and maximum deployment angles. The 2-dimensional geometry features output by the Matlab deployment model are fed into the Abaqus landing model in order to assess the landing performance with a single leg landing. The landing case has been chosen as a worst case scenario. 2 landing leg designs, with the same geometry, but one made of metallic parts and the other made of composite materials are compared in the landing model. In this model it was found that both designs pass the stress (or strain for the composite leg design) and G forces requirements when a shock absorber is attached to the leg and when they land on deformable Lunar like soil. On the other hand, the landing legs don’t pass G forces requirements when landing on stiff soil. Through the modelling, it was also possible to give preliminary design guidelines over the subsystems making the landing system. Therefore, system engineering practices have been used in order to initiate the design of some of the subsystems such as the landing damping, the leg joints and the locking mechanism to lock the legs at the desired angle and make sure they don’t retract. Simplifications have been made in order to model some of the landing and deployment aspects of the system for this thesis. Matters related to the adaptability and the detailed design of each of the subsystems have been left as future work. In the end, a comparison between the design developed in this thesis and the current Europa Lander Landing System design is made difficult by the lack of technical data on the latter related to mass. Nevertheless, a design that uses the decelerations of the engine module before touchdown is developed and the suggested material for the leg elements is CFRP. The design, though, didn’t pass mass fraction requirements when computing the mass of the extra subsystems. Even so, locations where potential gains through mass optimization can be performed for future studies have been pointed out.

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.

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.

Structural optimization for crashworthiness in composite structures has become an important topic of research in aerospace, attributable to its proven benefits to the occupants' safety in an aircraft during a crash event. This thesis provides a comprehensive investigation on LS-DYNA modeling and design optimization of composite square tubes for crashworthiness. The objective of this thesis is to construct a optimization framework specialized for the preliminary design of composite structures with a balance of performance and efficiency. Three primary components are involved: firstly, coupon-level simulations in LS-DYNA are performed to characterize the material properties for carbon/epoxy composite material IM7/8552; secondly, a square tube is modeled by a single-layer approach in three mainstream material cards (MAT-54, MAT-58 and MAT-262) with calibrated parameters for crush simulation in LS-DYNA. Meanwhile, detailed sensitivity analysis of influential parameters in different material models is also be performed to have a more comprehensive understanding of the complex failure mechanism. The simulation results indicate good correlation to the experiments in terms of energy absorption and maximum peak load, with high computational efficiency and low-cost calibration. Lastly, a two-stage single-objective optimization is performed, which incorporates the fiber orientation for each layer as design variables and design/manufacturing rules as constraints. Two surrogate models are created to formulate the mapping between input design variables and output crashworthiness metrics, including Deep Neural Networks (DNN) and Gradient Boosting Regression Trees (GBRT) ensemble. Followed by Mix-Discrete Particle Swarm Optimization (MDPSO) algorithm, the optimal set of design variables are obtained for each surrogate model. The first-stage optimization results demonstrate significant improvement in the crashworthiness performance compared with the baseline value, while the second-stage optimization results indicate an excellent transferability of the proposed optimization framework. This applicable, transferable, and data-driven optimization framework can be used in the aerospace industry regarding the crashworthy design and optimization of composite structures.

With increasing usage of fibre-reinforced composites for structural components across the automotive and aerospace industry, it became crucial to understand how this kind of materials behave when subjected to high loading rates, which are encountered in the case of crash events. The current work presents a study on the effect of the strain rate on the compressive mechanical properties of carbon-epoxy laminates. Rectangular and omega-profile cross-section specimens were tested on a servo-hydraulic testing machine at several impact velocities. For all cases, an increase in the mechanical compressive properties of the laminate was found with increasing rates of deformation. Another important aspect is the correct implementation of the strain rate dependent properties on the numerical models used to perform finite element simulations, as it can lead to more accurate impact simulations. A good agreement between experiments and simulations was obtained for both types of cross-section specimens considered.

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 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.

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.

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 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.

In recent years an exciting new era of space exploration has begun with the advent of the beyond Earth orbit CubeSat (BEOC) missions. However, utilizing the CubeSat platform for such missions gives rise to a series of significant technological challenges. Fortunately, the field of inflatable space structures offers some tantalizing solutions to these challenges. Unfortunately, at present the development of these structures is hampered by a technological gap in the required inflation systems. In order to fill this gap, this thesis proposes the development of inflation system optimized for such applications, based on the field of micropropulsion technology. In pursuit of this goal, the main research question shall address the design adjustments required in order to adapt current micropropulsion technology for this purpose. To answer this question the research objectives of the thesis are to first develop a preliminary design for an inflatable BEOC reflector and then, having generated inflation requirements based on this structure, design a custom micropropulsion based inflation system. This approach involves first identifying a BEOC inflatable reflector design whose continued development will play a key role in enabling and enhancing the development of BEOC missions and inflatable space structures in general. Then the structure is designed with particular focus on the key design considerations that impact the requirements of the inflation system. Following the design of the structure, the requirements for the inflation system are generated. Informed by these requirements, a trade off analysis identified a cold gas system as the most suitable micropropulsion system for inflation applications, as well as the inflation gas and a suitable inflation scheme. After this, the design of the inflation system is performed. It is found that in order to satisfy the inflation system requirements, apart from a unique design approach for the inflation nozzle, minimal design adjustments are required to cold gas micropropulsion technology for inflation purposes. This result demonstrates the feasibility of utilizing micropropulsion technology for the development of optimized inflation systems and provides a clear indication that such systems can play a key role in enhancing and enabling the development of BEOC inflatable structures.

Delamination is an important damage mechanism in composite materials. A correct prediction of its behaviour is of great importance in the design of aerospace composite structures. Nevertheless, when using simple numerical models which do not account for the delamination migrating to other interfaces or other intralaminar damage mechanisms, the accuracy of the numerical analyses is restricted by the inadequacy of the standardised experimental determination of the fracture toughness. This interlaminar property is usually measured using unidirectional specimens, which ignores the effect of the lay-up. Furthermore, the changing fracture toughness as the crack propagates, the so-called R-curve effect, is generally ignored as well by only employing the initiation values from the experiments. This leads to highly conservative results as the resistance to delamination is not correctly accounted for. The aim of this MSc Thesis was to contribute to the improvement of the blind simulation of delamination in co-cured composite aerospace structures. In particular, the experimental determination of the fracture toughness in coupons without a 0/0 interface to then be used as parameters in the simulation of one structural component: a single-stringer compression (SSC) specimen.Based on the lay-up formed by the skin and the stringer of the SSC specimen, various layups were defined for DCB and MMB testing. After evaluating them numerically, the fracture toughness was measured experimentally in unidirectional [012//012] and multidirectional [011, 45//-5, 011] coupons, in DCB and MMB for mode mixtures 20% and 50%. The R-curve effect of the multidirectional specimens was far greater than for the unidirectional ones, the steady-state fracture toughness was around three times higher, due to the crack migrating to other interfaces or propagating simultaneously in between different plies.The experimental results were used to improve the numerical analyses of the tests performed. A new approach was presented to model the measured R-curve effect by varying the fracture toughness values according to the distance to the crack tip. The improvements achieved were especially relevant for the multidirectional specimens. This approach and the R-curve from the multidirectional specimen were employed in the analysis of the SSC specimen. The results obtained emphasise the importance of accounting for the effect of the lay-up and the crack growth length on the fracture toughness. Furthermore, it was shown the potential and shortcomings of using an experimentally determined material law, accounting for delamination migration and other damage mechanisms, to improve delamination modelling.

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.