An analytical model for the onset and growth of delaminations induced by matrix cracks under static loading is proposed. The model introduces four distinct configurations of delaminations originating from crack tips, which are applicable to cross-ply and general symmetrical laminates subjected to tension or tension + shear.

The delamination onset method employs Kassapoglou & Socci’s analytical crack propagation model to describe matrix cracks within the laminate. This leads to closed-form expressions for the Strain Energy Release Rate associated with both crack propagation and the initiation of delaminations at crack tips. Both expressions are used to predict the delamination onset load and crack spacing. Predictions show excellent agreement with test results for cross-plies under
tension.

Regarding the delamination growth model, it assumes a constant distance between cracks during growth (crack saturation). Experimentally-obtained Mode II interlaminar fracture toughness equation and closed-form expressions for Strain Energy Release Rate for delamination growth are employed for predictions. The model overpredicts the initial delamination length but exhibits satisfactory agreement with test results for the delamination growth in cross-ply laminates under tension.

Combining all 3 models allows for a comprehensive prediction of crack propagation, delamination onset and delamination growth for general symmetrical laminates. This comprehensive approach enables the visualization of all relevant information in a single figure, providing a concise and informative representation of the damage processes. Moreover, the analytical model facilitates the construction of design curves to investigate delamination onset in detail.

Fibre Metal Laminates (FMLs) have been widely utilised in aircraft structures for their specific strength, durability and damage tolerance. Most aircraft structures, due to their size, require joining, and FMLs allow for a special type of joint called the splice, a unique but a lesser known joint.

The bulk of research into the splice joint has been conducted in the early 2000s for the A380 program, where a specific type of splice has been implemented, the overlap splice. While it is a proven design, it was expected that other splice designs could be as good if not better than the overlap splice. Further, the pre-existing design guidelines on the splices were not suited for detailed design that would target specific mechanical strengths of the joint, like durability and
damage tolerance. Therefore, this research focused on design iterations of the two splice types, the butt splice and the overlap splice, that would provide insights into splice design guidelines against fatigue and demonstrate which splice configuration is a superior joint. An extension
of this research focuses on identification of location and lifetime to damage initiation within the splice using analytical numerical methods, which was previously done on uninterrupted FMLs, but not on spliced FMLs.

Research comprises of a multi-disciplinary approach, combining Finite Element Method (FEM), numerical predictive modelling and experimental investigations that then also doubled as means of validation of the models built. Design iterations were based on the influence that the parameters, labeled a through e, within the splice joint have on the overall stress field within the joint and how that affects the joint’s fatigue life to damage initiation, Ni, a fatigue
performance indicator of the study. It has been discovered that changing the tolerances within the splices provides little influence on the stress field within the splice, which resulted in several iterations of the splice design approaching smaller and more lightweight joint alterations. The designs showed excellent durability characteristics when compared to the Limit of Validity (LOV) cycles set by the aviation authorities, such as EASA and FAA. The Finite Element (FE) model was able to accurately depict the stress field within the splice, validated by the experimental data through strain fields captured using Digital Image Correlation (DIC) technique, and consequently accurately pointed to the location of damage initiation, which in all cases was the outer-most overlap on the flush side of the joint. Along with the adapted predictive numerical model it was possible to predict the damage initiation lives in the spliced specimens with a blunt notch. The predictions in the samples without the notch resulted in far lesser agreement with experiments. This was likely caused by to the limitations of the model, which only took into account quasi-static loading without damage and is highly dependent on the reference data used.

It was concluded that splices can be designed smaller and lighter than previously done due to the tighter tolerances allowed within the splice. Overlaps of 5 mm in both the unnotched butt splice and the unnotched overlap splice, however, resulted in an alternative damage progression mode, specifically a complete delamination of the external overlap rather than metal fatigue cracking. This is attributed to the rising average shear stress in the adhesive. Regardless, the smallest and lightest iterations of the butt splice and the overlap splice with
5 mm overlaps and gaps between aluminium interruptions showed excellent durability and the design iterations were found to affect the damage initiation life very little, considering that fatigue damage initiation is a subject to scatter. It is hard to draw a concrete conclusion on the damage tolerance of the updated designs due to alternative damage modes and lack of research thereof. While both splice types were deemed to be successful in their role of a joining structure, the butt splice was concluded to be superior to the overlap splice when it came to fatigue performance, consistently developing visible damage later than the overlap splice. This gap in fatigue performance is expected to expand if thicker layers or larger number of layers are to be considered. This is because of the secondary bending which occurred in the joint, more so for the overlap splice than the butt splice judging from FEA results and experiments.

It is recommended that a more extensive FE model is developed using the model built in this research as a foundation. The current model could be expanded in several ways, such as simulations of damage progression, individual modelling of the fibre layers, fibre failure models, and a curing simulation. This will improve the validity of the methodology, specifically concerning the unnotched splice specimens, and improve the predictions of not only initiation lives, but also failure lives.

In composite structures, delamination damage is typically the most common failure mechanism. Accurately characterizing the delamination behaviour of composite laminates is therefore crucial for predicting the fatigue safe-life of such structures. To this end, double-cantilever beam (DCB) composite specimens are used to measure the interlaminar fracture toughness and delamination growth rate under cyclic Mode I loading.

In unidirectional (UD) composite laminates, delamination planes may exhibit fibre nesting, leading to the development of the fibre bridging effect during delamination growth. This effect, which resists delamination, significantly increases the apparent fracture toughness of the laminate. However, fibre bridging is usually insignificant in multidirectional (MD) laminates, where delamination occurs between plies with different fibre orientations. Nesting typically does not happen in MD laminates. As a result, MD laminates should not be designed using fatigue resistance data obtained from UD specimens without first accounting for the fibre bridging effect. Neglecting fibre bridging exclusion can result in an overestimation of delamination resistance, leading to unsafe failure predictions.

This research investigated methods to exclude the fibre bridging effect in cyclic Mode I experiments with UD composite specimens. Existing literature suggested different approaches to account for this effect, aiming to create a "zero-bridging" fatigue delamination resistance curve. The study examined methods such as cutting bridging fibres in-situ, constant-SERR experiments, specimen-specific extrapolation, and utilizing the Hartman-Schijve equation to describe fatigue delamination.

By examining different exclusion methods and understanding their limitations, this work contributed to enhancing the reliability of fatigue delamination predictions in composite specimens under laboratory conditions. This study compared methods to exclude the fibre bridging effect and assessed their merits in terms of ease of use, accuracy, and conservative predictions of delamination resistance. The results of this study suggest that a specimen-specific extrapolation method is a suitable approach to account for fibre bridging.

The concept of lightweight design is driving future aircraft to exploit all the available strength of materials and further reduce the weight of an aircraft leading to lower fuel consumption and more sustainable aviation. Current designs introduce rivets and bolts to join the structure and simultaneously create holes in the pristine material, which is not wanted due to the local stress concentration. If the design wants to exploit all the strength of the material, stress concentration should be avoided using suitable joining technologies that don't require holes in the structure. This is possible with the adhesive bonding technology. However, many factors, such as the effects of manufacturing and impact-induced damage are still not fully understood. Thus, the damage tolerance of the design cannot be guaranteed. This results in conservatory safety factors being prescribed in the design process, which possibly reduces the exploitation of all available material strength.

The occurrence of barely visible impact damage (BVID) in aircraft composite structures, especially in adhesive joints, is a serious issue that can jeopardise an aircraft's structural safety during operation. However, virtual testing and numerous numerical approaches allow high-fidelity simulation of damage initiation and propagation in adhesives and composite adherends to predict the residual strength of the bonding accurately. Although the development is fast, accurate impact simulation is still computationally very expensive and thus not applicable in industrial cases where behaviour needs to be analysed to assess the structures' design, maintenance and repair.

This thesis investigated two topics that allow robust and accurate simulation methods to assess the effect of manufacturing and impact damage on the residual strength of bonded joints. The first topic is related to the development of the composite material damage model, where the multiscale material model is created by the use of a representative volume element (RVE). The second topic of simulation methodology is the development of the quasi-static simulation approach for the damage tolerance assessment of bonded joints. Modelling approaches to simulate residual strength are investigated through the finite element modelling of three groups (i) pristine, (ii) artificially damaged during manufacturing, and (iii) impacted bonded joints. In the numerical models, an approach based on the observation of the fracture surface of single-lap joints in different geometry and layup configuration is proposed in which the damage assessment focuses on the bondline area up to the first 0° ply. For the modelling of the damage in joints, two modelling techniques were studied, first removing elements in the damaged area and second detachment of the elements. Utilising those techniques, simplified approaches to model damage resulting from the impact were studied, with the modelling of impact damage as a hole, where all through-thickness elements are deleted and as delamination, where interlaminar cohesive zone elements are deleted.

The developed multiscale material model allowed for an accurate representation of failure modes occurring in the composite material by the characterisation of elastic, damage and plasticity parameters of the fibre and matrix constituents in the homogenisation and inverse characterisation process. The results showed that the deletion of the elements can be used to represent defects and damage of different kinds in the composite bonded joints, both in the adhesive and adherend. The comparison of different representations of the impact damage in the single lap joint configuration revealed that the best prediction in terms of the ultimate load, failure mode and size of the model is obtained with the simplified representation as a single delamination positioned before the first 0° ply in the layup and in the studied adherend configuration that was between the first (45°) and second (0°) ply in the layup. The study ends with the conclusion that in different geometries of single lap joints and different damage types studied, a numerical analysis should focus on the region of the overlap edge and in the thickness direction from the bondline up to the first 0° ply in the lay-up. The results and numerical method developed during this thesis create a base for the further investigation of a variety of impact cases on bonded joints.

A structural supercapacitor (SSC) is a composite material that bears mechanical loads and stores energy, and its encapsulation method of the device is vital for scale-up. Here we present an SSC with an encapsulation interface consisting of spread-tow fabric electrodes and glass fabric encapsulators infused with epoxy resin. Through surface functionalisation of fibres, a high surface area electrode of 231.43 𝑚2/𝑔 was separated by a gel polymer electrolyte (GPE) separator and encapsulated in an epoxy polymer. To study the SSC’s performance, different GPEs are produced by adding lithium bis(trifluoromethanesulfonyl)imide (LiTFSI) into poly(vinylidene fluoride-co-hexafluoropropylene) (PVDF-HFP). The LiTFSI concentration is varied from 0% to 100% based on the weight of PVDF-HFP. The highest energy storage performance occurred at LiTFSI 60 wt% in the PVDF-HFP, with a cell-specific capacitance of 25.72 mF/g, cell-specific energy of 2.08 mWh/kg and cell-specific power of 25.04 mW/kg. With a flexural modulus of 3.39 GPa and a flexural strength of 52.07 MPa, this further develops high-performance SSCs in the future.

Fibre-reinforced laminated composites are constructed layer-by-layer, allowing their material properties to be easily tailored relative to their isotropic counterparts (metallic alloys). Moreover, a composite structure with variable stiffness (VS) can be made by spatially tailoring stiffness across a laminate. This permits better load distribution within a structure, efficiently using a given material, and allow for lighter construction. Typically, the design of VS structures follows a bi-level procedure: first, the structure’s stiffness distribution is optimised using lamination parameters (LPs), and second, the stacking sequences (SS) of differently oriented fibres are designed to achieve the desired stiffness distribution. However, the designs envisioned using LPs are only sometimes exactly matched by the designed SS. The shortcomings faced during this conversion are called the ’Inverse Problem’. Upon reviewing the existing techniques in literature to handle the inverse problem, it was understood that converting LPs into SS is challenging without incurring substantial computational costs or imposing significant restrictions on allowable fibre orientations in the design. Such restrictions tend to under-utilise the directional properties of the fibres. Thus, there was a need to explore and develop better stiffness design methods for laminated composites by bridging the gap between LPs and SS in a computationally efficient way.

In response to this challenge, a hierarchical design framework was proposed to handle the Inverse Problem. Diverging from conventional methods that directly design the SS and try to match a given set of LPs, this framework divides the problem into two distinct stages. Initially, the focus is to use the In-Plane LPs and determine the number of layers in each orientation within a laminate, also called the Fibre Angle Distribution (FAD). Subsequently, the FAD serves as an interim solution, and the SS can be designed by using the Out-of-Plane LPs along with it. This problem partitioning enhances computational efficiency and offers potential benefits in solving the Inverse Problem more effectively. Given the time constraints inherent to a master thesis, the primary undertaking of this study was to efficiently design the FAD while accommodating a wide range of possible fibre orientations. To address this, a novel method using the Fast-Fourier Transform (FFT) was developed to facilitate FAD design with ply angle multiples of 15° (or [∆15° ] = [0, ±15, ±30, ±45, ±60, ±75, 90]). Moreover, empirical guidelines for SS design, such as the Symmetry and Balancing rule, were incorporated into the design step.

The primary contribution of this report is introducing an FFT-based method to design FADs, a novel addition to the existing body of research. Upon extensive testing, it was shown that this implementation could convert LPs into multiple unique FAD solutions in less than 0.3 seconds on a regular office laptop, outperforming other methods in literature by at least ten times. Furthermore, the implementation was also used to demonstrate the benefits of designing laminated composites using [∆15°] over the conventional [∆45°] orientations (or [0, ±45, 90]). In light of the positive results, it is pointed out that the In-Plane LPs (and consequently the In-Plane stiffness) are known to be sensitive only to the FAD and not their SS. As such, the readers of this thesis are presented with a very computationally efficient approach for designing laminated composites for In-Plane Stiffness...

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.

Lugs are a specific type of joint with a semi-circular geometry around the pinhole utilized in the aerospace field. Its distinct geometry and ability to carry loads in its plane make it a very common choice for design engineers. Although this type of joint is widely used, there are no analytical equations that can predict the stresses around the hole. This research topic presents a semi-analytical approach to predict the in-plane stresses and failure modes of a composite lug under tensile pin loading. The methodology developed is based on stress functions for anisotropic beams implemented in stress equations for solid bodies in a polar coordinate system. The results of this approach are validated with the use of numerical simulation models, whose trustworthiness is verified by tensile tests. The outcome of this thesis is in-plane stress distribution graphs for every layer of the laminate around the hole and the in-plane failure mode of the lug. The results show a good accordance between the stresses predicted from the semi-analytical approach and the numerical simulations. Moreover, two lug designs with different geometric characteristics were tested in order to observe the influence of the geometry on the failure load. It was concluded that the geometry affects the maximum failure load but not the stress distribution, which is concurrent with the literature. Overall, the methodology presented in this research topic provided promising results. The findings of this thesis can have a great impact and aid engineers in estimating the stresses in a composite lug from the preliminary phase of a project.

Topology optimization (TO) has proven to be a capable design methodology for the realization of lightweight solutions that maximize structural stiffness or other design objectives. Due to its capability to be adapted to suit a wide range of objectives and constraints, both structural and non-structural, TO has been widely applied in all industries, including aerospace. The application of TO in secondary structures offers the scope for further weight savings and, therefore, this thesis investigates the employment of TO on a galley structure provided by Safran-Cabins. The galley structure is an essential equipment for the functioning of passenger aircrafts and hence, provides an ideal product, which once optimized could offer widespread weight saving.

The next generation of commercial aircraft require solutions that are lightweight and yet economic to manufacture. TO solutions are renowned for their complex architectures and frequent employment in conjunction with additive manufacturing. However, for products such as galleys, which are required to be manufactured in large quantities, complex manufacturing strategies are not economically feasible. Therefore, this thesis proposes a modularization strategy that can be used to augment the monolithic TO solution into an assembly of simpler and identical modules. The modular design strategy has proven to be a capable strategy to enhance manufacturability as well as reduce costs over the board and offers a unique opportunity to leverage the capabilities of TO for employment on a wider range of products.

To achieve the aforementioned goal, an image moment-based modularization strategy is investigated which treats the TO solutions as digital images and identifies positions of simple bar/beam modules within the image based on matching of the image moments. Through a detailed investigation, a fragment constrained bar matching strategy is developed. It is demonstrated that the proposed matching strategy is capable of identifying positions of bar/beam modules within TO solutions from the literature. Additionally, a post-processing strategy is developed to augment the obtained modules and their positions into simplified and assembled frame structures analogous to the underlying TO solution. The developed framework is employed on the topologically optimized galley, and its practical capabilities and limitations are identified, thus providing a foundational work to be further refined.