This thesis investigates matrix microcracking behaviour in thermoplastic composites at various temperatures, with a focus on Toray Cetex TC1225, a high-performance thermoplastic composite material reinforced with carbon fibres. The research examines the feasibility of using room temperature data to predict low-temperature performance, to minimize the need for extensive cryogenic testing. Both experimental and analytical methods are used to assess how mechanical properties and residual stresses influence microcracking. Comprehensive testing at room and low temperatures (-65°C) was conducted, with the resulting correlations analysed in detail. The findings of this study contribute to the development of more reliable cryogenic storage solutions, particularly

The demand for more efficient aircraft and new modes of transportation lead to a departure from the traditional tail and wing configuration. In order to operate an aircraft, it needs to be certified. Most Aircraft with a capacity to carry more than 10 passengers have to fulfil a variety of bird strike related certification standards. These standards define an impact velocity to design for, critical locations and tested areas however are chosen based on experience of previous aircraft. This is not possible if the design is radically different, as it is in the case of the Flying-V. The present thesis proposes a methodology to identify and rank all possible bird strike scenarios, based on geometry and flight path. The quantification of impact scenarios may offer a cost-effective way to assess and visualize vulnerabilities, ultimately reducing certification cost and time. Thereby allowing new concepts to be certified. The methodology synthesized in this report relies on decoupled analytical bird strike load models from the late 70s to quantify bird strike intensity. An algorithm has been developed to determine impact scenarios over the area of arbitrary computer aided design (CAD) geometries. Ray tracing allows for the exclusion of areas that can not be hit. The results are not sufficient to determine critical impact locations, as the damage is significantly influenced by the structural response. However, it is possible to quickly generate probable impact scenarios over large areas based on the flight path. Despite the discrepancy between predicted damage and impact intensity, intensity maps can indicate areas of interest for investigation.

This thesis investigates the effects of hygrothermal aging on the mode I and mode II fracture toughness of flax fiber-reinforced polymer composites (FFRP) under quasi-static (QS) and fatigue (F) loading conditions. A key motivation for this study is the potential of FFRP to replace synthetic fiber composites, as FFRP offers competitive mechanical properties while being biodegradable and less energy-intensive to produce. However, one of the main limitations of flax fibers is their susceptibility to environmental conditions such as temperature and humidity.

Delamination is a common failure mode in composites, and conducting fracture testing under mode I and mode II conditions is crucial for designing durable components. Double Cantilever Beam (DCB) and End-Loaded Split (ELS) specimens were manufactured for mode I and mode II tests, respectively. Subsequently, hygrothermal aging was simulated by subjecting the specimens to one or two cycles of humidification and drying at elevated temperatures within a climate chamber. Quasi-static testing was performed on unaged, 1-cycle aged, and 2-cycle aged specimens, while fatigue testing was conducted exclusively on unaged and 1-cycle aged specimens.

Testing resulted in significant plastic deformation of the specimens, this was attributed to their insufficient stiffness. This invalidated the assumption of Linear Elastic Fracture Mechanics (LEFM). To better capture these effects, the analysis was conducted using the J-integral, based on non-linear fracture mechanics. While the J-integral cannot account for all observed effects, it provides for a more realistic approximation for comparative evaluation of fracture toughness between aging states.

The results reveal that in mode I QS testing, the initiation fracture toughness on average improved by 19% after one aging cycle, with no further increase observed after a second cycle, while mode II QS fracture toughness was insensitive to aging. In mode I fatigue testing, a reduction in delamination growth resistance was observed after one aging cycle. Mode II fatigue testing exhibited substantial variability within aging states, making it challenging to determine the influence of aging, although a reduction in variability was noted after aging. The increase in QS initiation fracture toughness is likely due to the plasticization of fibers and matrix.

These results indicate that aging does not have a straightforward effect on fracture toughness, as its impact varies between modes and regions of crack growth. These findings provide valuable insights for the design of FFRP and other biofiber composites, contributing to the development of more sustainable materials.

This thesis investigates the design and analysis of cryogenic liquid hydrogen storage tanks in the context of aircraft retrofit. Hydrogen propulsion offers significant potential for achieving zero emissions in aviation, but its adoption introduces challenges related to safe and efficient storage.
The research focuses on comparing single-wall and double-wall tank architectures, assessing their ability to meet stringent operational, thermal, and structural performance requirements. In the preliminary assessment of a tank design’s viability, two key performance requirements are cruise time and dormancy time. A minimum cruise time of 20 minutes ensures the tank can support basic flight operations, while a dormancy time of 1 day ensures no hydrogen loss occurs if the aircraft remains stationary for an extended period, accounting for potential delays.
The methodology to calculate cruise time involves determining the maximum time the aircraft can remain in the cruise phase based on the inner tank dimensions, fill ratio, and mission profile. The dormancy time is the time required for the tank pressure to reach the venting pressure, at which point hydrogen must be released, and is calculated by implementing a thermodynamic model that simulates the tank's dynamic behavior over time, accounting for heat inflow from the external environment.
The evaluation of the single-wall tank reveals its simplicity and potential cost-effectiveness, but also exposes considerable limitations in terms of thermal insulation for the specific retrofit case study. This design approach is a viable option for larger-scale applications, where a lower surface area-to-volume ratio reduces heat transfer and, consequently, hydrogen boil-off. However, the compact dimensions of the tanks required for aircraft retrofitting present a significant challenge due to the inherently higher surface area-to-volume ratio, which leads to increased thermal losses and prevents the single-wall architecture from meeting the performance requirements imposed by this specific case study.
In contrast, the double-wall tank, equipped with a vacuum layer and multi-layer insulation (MLI), offers improved thermal performance. The heat transfer from the external environment is significantly reduced, allowing to preserve the cryogenic temperature of the hydrogen fuel. However, the added complexity introduces new challenges, particularly regarding the design of the inner vessel support system which must maintain the inner vessel’s position while accommodating thermal displacements and managing structural loads. Assessing the heat leakage budget for the support structure is the final critical step, as it determines the maximum allowable heat inflow through the support system, ensuring the tank meets its dormancy time requirements while allowing for design optimization.
The thesis develops a design methodology for the inner vessel support system, balancing the need for flexibility (to accommodate thermal contraction experienced during the first filling of the tank) with sufficient stiffness (to withstand operational loads, including emergency landing conditions). This approach involves selecting suitable materials and geometries that meet thermal requirements, while accurately determining the support structure’s stiffness properties. Different loading scenarios, such as normal operations and emergency landing conditions, are evaluated to analyze stresses and displacements in both the tank and support system. Adjustments to the design are made if stress or displacement exceed safe limits. The analysis reveals that optimizing the support structure is critical for the double-wall tank’s overall feasibility. While the double-wall design is technically viable and meets the thermal performance requirements, its success depends on further refinement of the support system to minimize heat leakage and ensure structural integrity.
The results of this study suggest that, although single-wall tanks are not suitable for this application, double-wall tanks offer a promising solution for retrofitting aircraft with cryogenic liquid hydrogen storage. Nonetheless, significant challenges remain, particularly in designing efficient support structures that can handle the operational demands without compromising thermal performance. Future work should focus on optimizing the support system design, exploring flexible materials, and considering additional factors such as sloshing loads to further improve tank reliability and performance.
In general, this thesis contributes to the development of a robust methodology for the preliminary design of cryogenic hydrogen storage tanks, providing a foundation for further advancements in hydrogen-powered aviation.

Aerospace manufacturers increasingly rely on composite materials for the most advanced aircraft, due to their superior performance and tailorability. As the adoption of such structures grows, so does the occurrence of various kind of damages throughout their service life. Therefore, there is the necessity to develop robust, reliable and repeatable procedures to fully restore the structural integrity of composite components. An additional challenge is posed by composite structures that include functional inserts: the functionality of such inserts needs to be re-instated on top of the integrity of the overall structure, while ensuring a seamless repair finish. This thesis considered two damage scenarios in a CFRP fuselage access panel whose edges are wrapped with a functional material, and defined the most suited repair procedure to tackle them.
The first scenario is represented by damages located in the CFRP structural part of the fuselage access panel. The scarf repair method was identified as the optimal one for this instance, research effort was therefore directed towards its improvement. Indeed, such technique does not currently allow to achieve a fully flush surface, as a mismatch between the repaired area and the undamaged one remains noticeable. Eliminating such unevenness is crucial for stealth and eventually aerodynamic reasons, and therefore needs to be investigated. Two repair configurations were implemented: the first one consists of re-milling the surface once the repair is completed, while the second one relies on a thinner repair patch that — once properly aligned — allows for a flush surface. The first option allowed for an improvement of over 80\%, reducing the surface unevenness from more than 1/2 of a millimeter to less than a 1/10, while ensuring a smooth, continuous surface finish. At the same time, it proved capable of meeting all the mechanical requirements, performing closely to a reference repair configuration in several tests. On the other hand, the second option only partially improved the surface finish, but fell short of the fatigue life requirement by a large margin and also showed a significantly poorer mechanical performance compared to the reference repair and the other configuration.
The second scenario is represented by small, cosmetic damages located in the functional edge of the fuselage access panel. It was identified that such damages are best addressed with repair procedures based on a filler compound. Guidelines to define such repair compounds were defined. Then, two repair procedures based on the use of these compounds were thought out: the first one is similar to conventional filling repair processes, while the second one relies on a bespoke tool to inject the repair compound. The former was implemented and produced promising results: it allowed to precisely restore the original profile and achieve a seamless surface finish. Damages affecting both the structural and functional parts were also briefly addressed, paving the way for future developments.

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.

A Helically wrapped structure is a hollow pipe structure that is made from a metal strip and the overlapping interfaces are adhesively joined. It is yet unknown how to determine the quasi-static strength and predict the fatigue life of this structure. A method was developed to determine whether either the adherend or the adhesive is critical to determine the strength and fatigue life. A machine was made to create the helically wrapped structure and tested for validation. It was found that the adherend is critical instead of the adhesive.

This research project investigates the potential use of ultrasonic welding as a tool to repair impact damage inside thermoplastic composites to restore the compressive strength. Compared to the already existing method, hot-pressing, ultrasonic welding could offer faster processing times, no need for dedicated tooling, smaller and lighter machinery and a smaller (heat) affected zone.
It was shown that ultrasonic welding is able to melt and re-consolidate delaminations, decreasing the damaged area. However, even when using the same welding parameters, the process showed large variations between the welds. These variations should be minimized for the process to be usable as a repair process.
Compression after impact (CAI) testing showed that welded repairs are able to fully restore the CAI strength. The re-consolidated material inside the damaged area suppresses local buckling, which normally causes premature failure. However, it was also shown that repairs with little re-consolidated material can decrease the CAI strength.

With a growing demand for longer blades in the wind turbine industry for higher rated power per turbine, a structurally sound blade-root connection is of commercial importance. Bushing connections have been a commercially favoured design in the past few years, replacing the commonly used T-bolt connection as the joining method of choice. The new design replaces the barrel nut in the T-bolt with an axial bushing that the bolt connects to and can be assembled with the laminate during the lay-up stage of the blade skin. It has been theorised that it can result in a reliable connection due to the elimination of laminate stress concentrations. However, literature outlining the performance of a blade-root connection with bushings is lacking in the current body of knowledge. While several patents for bushing designs exist, they don’t provide verifiable results on their efficacy due to trademark laws. The objective of this project is to design and conduct a numerical study of a blade-root connection with bushings with an aim to replace the T-bolt connection, along with providing evidence of the effect of various parameters on the structural performance of the blade root. The design is to be based on a Suzlon Energy-make blade with a pitch circle diameter of 3m and a blade length of 63m. The project has been planned in three phases: (1) Design of the root; (2) Validation of the design; (3) Comparative analysis. Modelling and FE analysis has been carried out in the ANSYS environment. Parameters of the bolted connection have been determined according to industry standards provided by VDI and GL. Design validation was conducted based on structural constraints; the key design constraint relevant to the blade-root as a sub-component of the wind turbine is the accumulated fatigue damage. For the final phase of the study, various parameters associated with the assembled root were identified and tested in iterations and their effect on the structural performance, weight, and cost of the assembly were studied. The results confirm the hypothesis of reduction of stress concentrations within the laminate; this eliminates several failure modes associated with composite laminates at the blade root. As is, the bushing connection can be considered a viable alternative to the T-bolt joint. Within the connection, higher absolute stresses and stress gradients were developed in the bolt joining the blade to the hub. Hence, this was the area of focus for fatigue damage evaluations. Conservative estimates of the accumulated damage show values well within the acceptable range. The connection has been designed keeping several concurrent variables in mind. Given the commercial applicability of the design, a rigid optmized design is not feasible due to unpredictable parameters like certification costs, total assembly times, and procurement costs. Therefore, an effort has been made to understand the effect of varying component parameters. The design lends itself to flexibility of dimensioning and material choice within the sub-components; parameters can be optimised according to cost, manufacturability, and performance requirements. While the base design configuration for the bushing connection is heavier than the T-bolt design, improved fatigue performance can be seen as a favourable trade-off.

Aviation is a vital lifeline for island communities. However, air travel is also a significant source of air pollution and therefore contributes to global warming. The rising sea levels, caused by global warming, heavily affect the livability of pacific islands, causing drink water shortages and crop failure. Sustainable aviation has to be a part of the climate change solution.

Sustainability has become an increasingly important issue, and several different governments around the world have been working towards more environmentally friendly approaches throughout different industries. This has led to measures such as the European Green Deal, which aims to make Europe climate neutral by 2050. For the aircraft industry, however, this goal creates a so-called circular causality problem. This is because there may be limited investment in increasing production of alternative energy sources due to the limited availability of aircraft that use them. On the other hand airlines may hesitate due to the limited availability of fuel to buy such aircraft. In order to solve this problem, the carbon neutral ready aircraft has been proposed. The carbon neutral ready aircraft is designed such that it initially is powered by fossil fuels and can then be converted to be powered by a carbon neutral energy source. The carbon neutral energy source that is chosen for this aircraft design is synthetic kerosene. The carbon neutral aircraft has a high wing configuration with a high aspect ratio wing. To cope with the large span of the aircraft it was decided to give the aircraft the ability of the wing tips to be folded up. Furthermore, the wing has support struts which connect the wing to the lower part of the fuselage. The propulsion system of the aircraft has two novel features: two wing mounted ultra high bypass ratio turbofan engines and an electrically powered ducted fan, which ingests the boundary layer at the aft of the fuselage. An additional five-gear configuration was chosen for the landing gear to provide the aircraft with stable ground operations while not creating a need for a fairing which interferes with the boundary layer being ingested by the aft ducted fan. Carrying out the design of this aircraft shows that the aircraft is financially feasible and performs as well as the baseline A320 aircraft in terms of payload and range, while allowing sustainability goals such as the European Grean Deal to be met. The aircraft has a 17 % emission reduction compared to the A320neo, while still employing fossil fuel based kerosene. Furthermore, at least 90% by mass of the primary structure of the aircraft is recyclable. From the recommendations however, it is clear that a lot still needs to be done before the carbon neutral aircraft can enter into service in 2030.

The Mars One initiative is determined to put the first humans on the surface of Mars. Furthermore,
it aims to create a permanent human colony on Mars by providing a one-way transportation
service to the red planet as early as 2023.