The established theory of diagonal tension (i.e., post-buckling of skin-stiffened structures under shear load) for curved skin is based on experiments conducted on stiffened cylindrical shells at NACA in the 1940s-50s. However, the experiments were unable to capture all aspects of the phenomenon, leaving knowledge gaps in the understanding of the local distribution of the compressive loads generated on the stiffener, which were filled using the theory of diagonal tension for flat skins. Furthermore, due to the limitations of the time, the secondary bending effects on the stringers were not captured well.

This MSc thesis attempted to fill this gap by performing virtual tests using non-linear finite element modelling, which is enabled by the significant improvements in computational hardware of today. A virtual testing framework is developed based on the stiffened cylinders tested at NACA using the ABAQUS finite element modelling software. The skin and the stringers are modelled using continuum shell elements and the rivets are modelled using point-based fasteners. Furthermore, the generation of the models were parameterised using python scripting with provisions to model imperfections on the cylinders in the shape of the buckling eigenmodes. To validate the modelling framework, FE models of 8 cylinders were created and the results were compared with experimental data. Satisfactory correlation was found between the virtual testing results and the experimental results given the under-reported experimental data and the unknown imperfections in the original test

A custom post-processing method was developed that was able to capture the axial and bending loads that were generated in the stringers, which were used to develop a more detailed understanding of the load-transfer mechanism under diagonal tension. The results were also compared with the semi-empirical design method and the theory of diagonal tension, and significant discrepancies were found pertaining to the local variation of the axial loads and the secondary bending of the stiffeners.

Fibre-reinforced polymer (FRP) composites have become indispensable in aerospace engineering over the past two decades, driving the need for efficient and standardized testing protocols to certify their reliability. However, no standardized protocol currently exists for fatigue testing of FRPs, particularly due to fibre bridging—a phenomenon prominent in testing but rarely encountered in real-world applications. This study investigates existing fatigue data processing techniques, with a primary focus on a regression-based approach that could standardize fatigue data analysis. Current fatigue delamination characterization methods often rely on single-parameter empirical models, which struggle to capture the complex interaction between cyclic and monotonic load components, expressed as ∆√G and Gmax, respectively. This study demonstrates that the regression method’s zero-bridging technique effectively incorporates both parameters, offering a comprehensive view of delamination growth by isolating it from fibre bridging effects. This approach suggests a shift away from the traditional 2D analysis to a 3D framework, which considers both Gmax and ∆√G, enabling a more accurate depiction of delamination behaviour across various stress ratios. Notably, the results reveal that zero-bridging data align on a common plane under consistent stress ratios, shifting as stress ratios increase. Additionally, fibre orientation influences data clustering, with similar orientations exhibiting stronger convergence than dissimilar ones. Comparison of this regression-based method with the modified Paris law reveals significant discrepancies in its current implementation, suggesting alternative approaches to address these limitations. This study also highlights the impact of data size and selection on the model behaviour, stressing their importance in accurate model representation. This research validates the regression method as a promising candidate for standardizing fatigue data processing, improving the precision and reliability of post-test analysis for fibre-reinforced polymer composites.

This thesis explores the potential of sustainable resins in composite manufacturing, focusing on bio-based benzoxazine resins and bio-based epoxy derived from brown algae. The research investigates the feasibility of integrating these sustainable alternatives into composite structures. Bio-based benzoxazine resins were synthesized in-house, utilizing m-guaiacol, furfurylamine, sesamol, paraformaldehyde, Jeffamine T403, and DDS, while a commercial bio-based epoxy, phloroglucinol triepoxy (PHTE), was employed. The study involves comprehensive material analysis and testing, including an examination of chemical properties and mechanical performance. The findings reveal contrasting outcomes for the two types of bio-based resins. Bio-based benzoxazine resins exhibited challenges, as curing was not achieved, preventing their combination with composites. In contrast, phloroglucinol triepoxy, the bio-based epoxy from brown algae, emerged as a promising candidate for sustainable composite investigations. This resin demonstrated excellent mechanical properties, although its high viscosity and reactivity posed challenges for traditional manual manufacturing techniques. Furthermore, resin characterization unveiled a high glass transition temperature (Tg) system, especially noteworthy given the use of an aliphatic curing agent and monofunctional reactive diluent that lowers crosslinking density. This opens the door to its application in composites across various industries.

Leading Edge Erosion is a major concern for the wind energy industry as it affects the aerodynamic efficiency of wind turbine blades and, therefore, the power generated. Hail impact on the leading edge could significantly damage the protective layer and the composite substrate. The focus of this research study is to investigate the effect of kinetic energy on the damage mode in gelcoated glass fibre reinforced composite in terms of the velocity and sizes of the hailstone during hail impact. The gelcoated glass fibre composite samples were tested using a gas cannon using parameters based on conditions experienced in real life but with slight modifications to enable accelerated testing. The gelcoated samples were impacted at various velocities for different hailstone sizes, 15mm hailstone (140 m/s, 150 m/s and 160 m/s), 18mm hailstone (100 m/s, 110 m/s, 120 m/s and 140 m/s) and 20mm hailstone (90 m/s) to determine the Failure Threshold Energy (FTE) for each hailstone size considered. This threshold was determined to be the limit below which no damage would be initiated. The kinetic energy of each hailstone is estimated from the mass of the hailstone, and impact velocity, which is measured using a high-speed camera. Using non-contact profilometry (optical microscopy), cross-sectional damage analysis, and ultrasonic C-scan, the damage on gelcoated fibre-reinforced polymer composite samples were analysed.
The coatings did not exhibit any damage or signs of delamination in the coating-substrate interface. However, the substrate showed signs of impact damage. The damage mode was highly influenced by the kinetic energy imparted by hail impact. The damage mode of matrix cracks in the transverse and longitudinal directions was observed in the substrate for most of the impact parameters. The damage mode remained the same for the different hailstone sizes. The impact tests confirmed that a Failure Threshold Energy existed for each hailstone size. It was observed that FTE was increasing with the hailstone size, and it could be stated that multiple impacts had no effect on damage initiation for a limited number of impacts.

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.

Microbes, and the biofilms they form on surfaces, generally have a negative impact on performance. Well­-known examples are the increase of drag in ships and pipes, and the accelerated corrosion of metallic structures known as MIC. Among microbes, diatoms form a large unicellular algae group known to play a relevant role in the first stages of biofilm formation. Diatoms are well-­known for their species­-dependent silica exoskeletons, but they also produce, as other microbes, extracellular polymeric substances (EPSs) that may offer opportunities for the development of novel bio­based surface treatments. In this work we studied the interaction between a marine diatom species named Cylindrotheca fusiformis and aerospace aluminum alloys as a first step towards novel surface protection treatments of aircraft structures.

Two routes were followed to study the interaction of C. fusiformis with aluminum alloys after studying diatom population growth. The first route focused on the biofilm formation on aluminum 99.5%, AA20204-­T3 and AA7075­-T6 to study the effect of surface composition. The effect of the surface preparation (degreased vs. polished) was then studied with AA2024­T3 substrates. The second route focused on the study of diatoms’ motility on aluminum 99.5% and AA2024-­T3 with various surface pretreatments. In this work, a number of characterization techniques such as FTIR, SEM/EDS, Raman spectroscopy and image correlation techniques were used.

It was found that C. fusiformis forms well­-adherent biofilms on all aluminum substrates, independently of the surface composition and surface pre­treatment. Nevertheless, the biofilms appeared to be most homogeneous on AA7075­T6, while the EPS layer was more homogeneous on aluminum 99.5%
and AA2024­T3. The motility studies showed that C. fusiformis lowers its surface motility rapidly when on aluminum 99.5% until cells stop being motile. On AA2024­T3, cells maintain their motility for longer periods of time. The motility dependency on surface chemistry is hypothesized to come from speciation dependent aluminum toxicity. Post­mortem chemical analysis of the exposed surfaces showed traces of organic material in the form of proteins and carbohydrates attributed to displacement trails. The results confirm that monocultures of C. fusiformis are able to generate biofilms on aerospace aluminum alloys and release water­ insoluble organic matter on the surface. The promising results here obtained pave the way for more dedicated research to understand the relationships between surface chemistry and topology, and diatom motility and biofilm formation and composition.

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.

The demand for efficient lightweight structures grows rapidly in the aerospace sector. Topology optimization, introduced in the late 80s, is a method capable of producing such structures and has been mainly studied for isotropic materials. On the other hand, the performance of composite structures, which are already widely employed in the aerospace industry, can be now improved due to the latest advances in manufacturing. Automated Fiber Placement has paved the way for further exploiting the capabilities of composites and led to the introduction of an optimization method, namely three-step variable stiffness design, that alters the fiber orientation within each ply, leading to a variable stiffness laminate. In this thesis, a staggered optimization method that simultaneously improves the material and the structural performance of balanced and symmetric laminated composite structures under planar loading is developed by coupling the two aforementioned methods. A finite element code is developed based on the equations for a static and linearly elastic problem. For the topology optimization problem, the nodal density values are used as design parameters and the Solid Isotropic Material with Penalization approach is chosen, along with a density filtering. The first step of the variable stiffness design method is implemented, using the nodal lamination parameters as design variables. The coupling of the two optimization techniques is performed through a staggered optimization, where both techniques are implemented successively, at each iteration. A gradient-based scheme is used for both topology optimization and variable stiffness design and the steepest descent method is implemented for the design update. The sensitivity analysis is performed based on the continuous adjoint method. Three different examples are demonstrated in this thesis; the case of a flat composite plate, a lug and an aircraft chair bracket. The whole optimization procedure, including pre- and post-processing, is carried out using an algorithm developed in MATLAB. The meshes are externally created and imported in the code. Convergence studies and a comparison with indicative examples from the literature are performed in order to verify the obtained results. A parametric analysis of the plate studies the effects of the different topology optimization parameters. Finally, comparative analyses are executed for the three aforementioned designs investigating both the structural and the computational efficiency of the results obtained with each of the three optimization methods, i.e. topology optimization, variable stiffness design and staggered optimization.

Welding is a joining method for thermoplastic composites (TPCs) that offers multiple advantages over the more traditional methods of mechanical fastening and adhesive bonding. A particularly promising welding technique is ultrasonic welding, which features very short process times as a result of the high heating rates that can be achieved. This spot welding technique is hypothesized to have potential for improved damage tolerance compared to more commonly used continuous welding techniques: in a multi-spot welded joint, evolving damage will need to re-initiate in subsequent spots. The fact that damage initiation will need to occur multiple times might delay overall damage evolution through the joint compared to a continuous welded joint, where damage initiation needs to occur only once. This work is a first exploratory step into the domain of fatigue of multi-spot welded joints in TPCs. Existing research on the fatigue behavior of four-spot welded steel joints in various layouts served as the main reference throughout this research: its methodology was transferred to four-spot welded joints in TPCs. By comparing fatigue behavior across both materials, it was evaluated to what extent existing knowledge and design rules for steel could potentially be transferred to TPCs. Differences were observed in the results obtained for TPC and steel joints. Most notably, in steel joints the dominant failure mode was seen to change from spot fracture to sheet fracture at higher fatigue lives. In TPCs, joints consistently showed spot fracture across all load levels. A different interrelation between layout performances was seen in the steel and thermoplastic composite joints, assumed to be a result of localized material strengthening in the steel joints from interference of adjacent heat-affected zones. These results indicate that existing knowledge on multi-spot welded joints in steel cannot be readily transferred to TPCs, as failure modes and material mechanisms may differ. It was discovered that, when one spot failed prematurely as a result of existing damage in the joint, the remaining layout no longer seemed to have an effect on fatigue life performance. This was attributed to asymmetry in the remaining joint layout, meaning one spot would always become a preferred location for damage initiation and subsequent evolution. Therefore, subsequent damage evolution would only be restricted by a single spot up to the point where the shear strength of the joint was exceeded.

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.

Induction welding is an effective technique for joining unidirectional carbon fiber reinforced thermoplastic composites and L-joints can be produced through quick and cost-effective processing steps. However, due to high localized stresses in the skin-stiffener interface, these L-joints are often avoided in primary aircraft structures. Also, no international testing standards have been developed for testing of such joints.

A method was developed for the implementation of a neat thermoplastic resin fillet between the L-joint skin and stiffener web using the induction welding process in an attempt to remove the high stress concentration at this location. A 35.4% increase in quasi-static pull-off strength was measured with a weight penalty of less than 0.5%. This result was compared with a similar autoclave co-consolidated joint, which showed an 80.9% improvement. An ANSYS Parametric Design Language finite element model was developed based on the virtual crack closure technique and it showed that the joint pull-off performance is strongly dependent on geometric parameters such as the skin and stiffener thickness. Also, a new test setup was developed, which reduced internal stresses created by the setup compared to those commonly used in literature.

By further improving the method through which the fillet is joined to the induction welded L-joint, a performance increase similar to that of the co-consolidated joint should be achievable. Test results have shown that the use of this type of fillet can lead to the skin-stiffener interface no longer being the critical failure point for realistic joints in primary aircraft structures.

A numerical fatigue delamination model is validated by means of experimental tests. Delamination is initiated on the skin-stiffener interface which, under cyclic compressive loading and in the post-buckling regime, grows at a decreasing rate. A cohesive zone formulation is used to model the static and fatigue delamination growth, good agreement is found in terms of delamination shape. Further investigation is required, using improved material inputs, for the model to be validated in terms of delamination magnitude. The numerical model shows promising characteristics, which may be employed for future damage tolerant composite structural designs.

New methods for aircraft movables introduce oblique- and out-of-plane loads on lug joints. With the introduction of these loads on lugs the need arises for a more detailed fatigue life prediction method.

Therefore, the thesis research focused on creating a methodology for predicting the fatigue life for lug joints subjected to combined in- and out-of-plane loading in co-operation with GKN Fokker Aerostructures. Current prediction methods are all based on axially loaded lugs. In concept, this method applies the Larsson relation.

From Finite Element Analyses (FEA) of different in- and out-of-plane loaded lugs, the influence of eccentricity, taper angle, load angle, and lateral loads on the Larsson relation became apparent. Resulting in the addition of correction factors K_ecc, K_α,and K_L in Larsson. Although the method is verified, experimental test have to be performed for validation. Recommended is the expansion of the out-of-plane FEA and the inclusion of transverse loaded lugs.