Society increasingly demands that ships become more sustainable and quieter, given their significant share of global fuel consumption and green-house gas emissions. Fiber-reinforced composite marine propellers can contribute meaningfully to these goals. As a promising alternative to the conventional rigid metallic propellers, flexible composite propellers can offer improved underwater radiated noise (URN) and propulsion efficiency. Furthermore, manufacturing propellers from fibre-reinforced composite materials makes them lighter and reduces their electromagnetic signature.

Despite the significant potential of composite propellers for marine propulsion systems, uncertainties in their fatigue behaviour have so far hindered their wide-spread adoption. These uncertainties can arise from imperfections during the manufacturing process, operational conditions different than the ones considered in the design, coating deterioration leading to water ingress, impact events, and more. Such factors can have significant impact on the lifetime of the propeller, which is typically expected to endure billions of cycles. Structural health monitoring (SHM) has the potential to mitigate this issue by real-time recording and assessing the structural response and integrity of the propeller. Such an SHM system should neither affect the propeller performance nor its load-bearing capacity. In addition to providing insights into the current structural integrity of the propeller, an SHM system may also enable enhanced estimation of the remaining lifetime, thereby minimizing the risk of unexpected failures and downtime.

This thesis investigates the feasibility of developing composite marine propellers with an embedded SHM system based on piezoelectric sensors. These sensors are capable of performing strain monitoring (with application in load/response estimation) and acoustic emission monitoring (with application in damage identification). Three main topics have been studied; (i) the feasibility of measuring dynamic strains in the propeller blade using the embedded sensors, (ii) the effect of embedding piezoelectric sensors on the structural integrity, and (iii) the feasibility of measuring and assessing damage-induced acoustic emissions using the embedded sensors. An analysis framework has been proposed for the identification, classification, and localisation of acoustic emissions in thick composite structures…

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.

Fiber reinforced composites have been increasingly used in the aerospace and automotive industry, due to their potential advantages for designing flexible, strong and lightweight structures. More recently they are also being considered for the manufacturing of marine propellers. Since they have the potential to lower the weight, lower the maintenance costs, increase the efficiency at off-design conditions, improve cavitation inception speed and minimize acoustic signatures. Exploiting the full potential of composite propellers however requires that they need to be cost-effective and to do so it is required that the lifetime of the blade is sufficient. To determine the lifetime of the blade it is
crucial to determine how the composite blade will respond to a wide variety of environmental and loading conditions that it will experience over its life. Basically one needs to determine what effects fatigue will have on the material properties of the blade. The main objective of this thesis is to estimate the fatigue lifetime of composite marine propellers subjected to a pressure distribution determined by a Fluid-Structure Interaction
(FSI)model with the use of a progressive damage model integrated into a Finite
Element Model (FEM). The fluid structure interaction model to determine the pressure distribution has already been created Maljaars (2019). Currently the mostly used and most reliable approach is to use an experimental approach in order to estimate the fatigue life. However, for custom designs such as marine propellers this is a long and costly process. Modelling fatigue the fatigue performance in an earlier stage of the design cycle will result in significant cost and time savings.

The expected fatigue lifetime of the composite marine propeller is concluded to be at least 1012 cycles under the considered loading condition. If the propeller would rotate at 600 rpm for 24 hours per day this would translate to over 3000 years. The reason behind this large number is that the combination of the applied pressure load and the strength of the carbon fiber propeller is such that the stresses in each ply are very low. These low stresses create almost no damage throughout the propeller blade even for ultra high cycles.

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.

The use of Grid-stiffened structures has been increasing in the space industry over the last decade. Thanks to it is high specific mechanical properties and excellent damage tolerance characteristics, researchers also started looking into the potential of using them for aeronautical applications. However, not much has been done to investigate the reduction in strength after barely visible impact damage and visible impact damage, and how much energy is needed to cause such damage, which is an important requirement for aircraft certification. In this project, a grid-stiffened replica of an existing reference cargo drone is created, and the design is tested to check its compliance with the requirements. This involved modelling of the design, manufacturing of test coupons, impact and mechanical tests, and model validation. The results showed a possible weight reduction, with no major reduction in strength observed with impacting using the cut-off energy level set by the manufacturer.

A building block pyramid is designed used to evaluate a composite aircraft structure, combining tests and analyses, using stiffened panels, single-stringer specimens and coupons. Failure in these structures can be caused by postbuckling-induced skin-stringer separation, which is complex, involving matrix cracks, fibre bridges, and delamination migrations. Standardized tests covering postbuckling-induced skin-stringer separation are lacking. First, a material characterization is performed on the coupon level. These properties are applied to a stiffened panel model to identify critical postbuckling regions. Then a single-stringer specimen is designed that combines the material complexities of the coupon level and the geometrical complexities of the panel by mimicking the postbuckling shape. Specifically a seven-point bending configuration to study bending-induced separation and a four-point twisting configuration for skin twisting-induced separation. The guidelines for modelling and testing can assist in standardising these test methods.
Using composite materials in aircraft structures can reduce weight compared to conventional metals. However, utilizing more of the material's load-carrying capabilities can further reduce weight.

Fiber Reinforced Composite (FRC) materials are gaining great popularity in marine structures because of their excellent strength-to-weight ratio, low density, and provided freedom in the design process. However, the use of FRC materials comes along with relatively large uncertainties in material properties and structural integrity after manufacturing and during its use. In this research a new in-situ non-destructive stiffness assessment methodology is proposed. This methodology is based on a coupling principle between the laminate structural stiffness properties and the ultrasonic guided wave characteristics of FRC materials. In the methodology, a range of possible stiffness properties is defined based on the structural information available for a structure of interest. The average relation between this stiffness range of interest and corresponding wave characteristics is described using a set of coupling coefficients which are determined using numerical simulations. For this, a batch of reference laminates is constructed that covers the entire stiffness range of interest. Input for the system are the group velocities of the zeroth-order symmetric and antisymmetric guided wave modes, measured on the structure of interest. The potential of the proposed methodology is evaluated using a numerical feasibility study based on numerical simulations. Good results were obtained for different test scenarios, varying in the amount of structural information available on the structure of interest. Thereafter, the in-situ application of the methodology has been examined in an experimental setup. Good measurement and analysis times were achieved by using a compact measuring device that is capable of recording the wave signal in five directions simultaneously. A reliable accuracy assessment of the in-situ application of the methodology was, however, difficult to obtain due to the lack of reliable reference studies. Therefore, in future research reliable reference information should be gathered. Additionally, a next step for future research should focus on decreasing the amount of information available on the structure of interest.

This study is an attempt to develop a preliminary composite fuselage sizing tool for the decision making process utilizing analytical structural analysis equations and a bottom-up cost approach to estimate weight and cost. Previous research by several authors and companies is reviewed in order to define the knowledge gaps that the project is going to fill. The objectives of this project are divided into two directions. The first one is the weight and cost fuselage design, and the second, is the weight and cost optimization process. After examining the existing literature both from academia and industry a different approach is proposed for each sector. In the structural design an analytical based model is proposed to assess the components integrity while cost equations for fully learned manufacturing processes are used based on the Advanced Composites Cost Estimation Manual (ACCEM). An experience-driven optimization algorithm is proposed afterwards, taking into account manufacturing considerations and design rules of thumb. By combining these tools, a general tool for fuselage design that is completely suited for the preliminary design phase is proposed. Τhe trade-off studies can provide important knowledge on the effect of different decisions and establish new design guidelines. The study concludes that an increase of the number of stiffeners and frames, in a fuselage structure, leads to a general decrease in weight and an increase in manufacturing cost. The addition of more stringers or frames after a certain level leads to an increase in weight, while the cost follows the same trend. The manufacturing constraints and the design rules-of-thumb applied lead to an increase in structural weight. Finally, the comparison between skin-stiffened and sandwich designs, shows that the weight saving potential is comparable for both configurations, and the cost saving potential is higher for the sandwich design.

The Paris agreement’s mechanisms for 2050 represent a challenge for the world to achieve Net-zero emission and climate resilience. Retrofitting the existing stock is a critical step in every nation to achieve these goals. The building sector has a significant role in the Net-zero emission transition. The ambitions for retrofitting the Netherlands´ by 2030 and 2050 bring the need to explore new products taking sustainability and circular economy as the basis of design. This research explores the application of a “cradle to cradle” design approach to redesign a sandwich panel currently used as a renovation strategy to wrap existing dwellings in The Netherlands. The research in performed from the perspective of a façade company in the national market.

The current product manufactured by the company is not studied with the end of its service life in mind and is designed mainly with fossil fuels related products. Also, the time component is detached from the product, and scenarios where the materials are “processed and disposed” or “mined and reused” are not considered. The research explores three different façade concepts that contrast with a traditional linear production based mainly on fossil fuels. The analysis brings a set of 24 options, each with three circularity scenarios. The conclusions reveal that the environmental impacts and success of a “cradle to cradle” design strategy has a close relationship with the number of years the existing dwellings will be used. By reusing the existing dwellings for prolonged times (50 and 100 years), the best option for the company is to develop a biobased sandwich panel relying on renewables and materials with low environmental impacts but as an efficient “cradle to grave” strategy. However, for a shorter span of usage in the existing stock (25 years), the best option is a “cradle to cradle” strategy where the resources are taken back to the technical cycle combined with reduced usage of materials for the cladding system.

Some of the technical recommendations suggested are to test the biobased panel for a mechanical test. Afterward, develop the construction details for connections in foundation, windows, and doors to finally build a 1:1 mock-up to be tested for meteorological degradations and durability. Also, further analysis is needed for a financial case for the scenarios where materials are used after a first cycle. Finally, further research is needed to develop fully biobased matrixes to biocomposite fully biodegradable, allowing them to get back into a biological cycle.

Straight fiber variable stiffness laminates use multiple overlapping patches to tailor the fiber orientations in particular regions of the laminate. A novel, automated framework for the design of manufacturable laminate has been developed using different stages. In the first and second stage an optimal stiffness distribution is determined using lamination parameters and converted to stacking sequences on a per region basis. In the third and fourth stages cellular automata and network theory are used to introduce cohesion between regions and comply with current manufacturing standards. Plates in uniaxial compression and cylinders in bending show an increase in buckling performance of 42% and 30% respectively. Such improvements do allow weight savings between 7-11% for these types of structures. When the additional cost of manufacturing and the cost-to-weight ratio for airlines are considered, these laminates already have the potential to significantly reduce the cost of air travel.

The COVID-19 pandemic had a significant impact on the aviation industry with more than 60% reduction in the passenger traffic in the year 2020 compared to year 2019. The passenger traffic which is now expected to reach the 2019 level only around the year 2024 will still continue to grow but at a lower pace compared to the pre pandemic levels. The environmental issues which were a major concern for the aviation industry before pandemic will still be relevant. The objective of achieving an environmentally friendly zero emission aircraft will not be met only with alternative fuels and propulsion concepts but also require advanced material technologies and novel designs. A promising technology having the potential to improve the performance of an aircraft by improving the structural efficiency is the application of aeroelastic tailoring with the help of composite materials. However, incorporating aeroelastic tailoring with composite materials in the design process is not a trivial task. In the traditional design process, knowledge about the design increases, while the design freedom decreases as one goes from conceptual to preliminary and finally to the detailed design. For conventional designs, the lack of knowledge during the initial stages is compensated through empirical knowledge. However, the lack of such empirical knowledge for novel design and advanced technology, results in the need for increased physics-based knowledge during the initial design process. In the research presented in this dissertation, the focus was on increasing knowledge in the early stages of the aeroelastic design process of a composite wing. As current state of art in aeroelastic tailoring does not include critical gust and fatigue loads, this thesis is focused on including critical gust loads and fatigue loading requirements in the preliminary aeroelastic optimization framework…

Composite lattice structures (CLS) offer high performance and demonstrate significant mass savings in space structure applications. Local modifications of regular lattice designs have the potential to further improve the lattice performance. This research explored the micro-structural quality features of CLSs manufactured with pre-preg fibre-placement, and how quality is impacted by lattice modifications.
Modifications to a regular lattice for a representative case of an attachment point load are identified using a topology optimisation tool. Three lattice modifications techniques were identified: rib width variations, rib angle variations and additional ribs. A sample lattice with modifications was manufactured and evaluated by C-scan, CT techniques and micro-sections. An explanation of quality features in CLS is described and quantified using a presented characterisation model based on node transition waviness.
Using the presented explanation, a quantitative quality model was developed to relate lattice design geometry to manufactured quality for this process. The model was used to design and manufacture a second panel with improved implementation of lattice modifications. The quality model was improved and shown to explain more than 75% of all observed quality variation in the two panels with a linear regression. Future work and limitations the of the quality model are discussed.

An analytical model for stiffness degradation of composite laminates with damage under static or cyclic loading is proposed. For static loading, two different modelling approaches have been created. Both account for shear-induced microscopic matrix damage and matrix cracking within each ply orientation of a laminate. In one approach the residual transverse tensile and shear moduli of a ply are determined by equating the energy density of the undamaged ply to that of the ply with damage, for a same applied load. Predictions excellent agreement with test results for cross-ply laminates, while stiffness degradation is for shear dominated layups. The other static approach applies energy equivalence only to the transverse tensile behavior of a ply. For shear, it predicts the residual shear modulus by accounting for the creation of permanent shear strains. Agreement with test results is excellent for cross-ply laminates, and excellent to good for shear dominated layups. The physical foundation behind this model is not as rigorous as the previous one, and as such needs more work. The proposed fatigue model assumes the matrix strength of a ply to be randomly distributed. Kassapoglou’s residual strength model for fatigue loading is used to express fatigue life as a function of matrix strength. The static model is then used to estimate stiffness degradation as a function the same matrix strength. Connecting the two, stiffness degradation as a function of fatigue cycles is obtained. Validation was performed on cross-ply laminates and agreement with results is excellent.

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.

Fibre-reinforced composites are increasingly being used in maritime structures. Piezoelectric sensors can be used to monitor both loads acting on a fibre-reinforced composite structure and elastic waves emitted from damage initiation within the structure. In certain marine applications, due to a harsh environment or for hydrodynamic reasons, piezoelectric sensors cannot be placed on the surfaces of the structure. Fortunately fibre-reinforced composite materials allow piezoelectric sensors to be embedded inside the structure. The focus in this research is on embedded piezoelectric sensor design, performance and behaviour. Carbon fibre beam specimens with embedded piezoelectric sensors have been manufactured and subject to low-frequency bending, high-frequency elastic wave emission, electrical excitation and monotonic bending to failure. The low-frequency bending test has given insight in the effects of sensor discharge, that is, decreasing and distorting the sensor response. Also it has been used to check consistency between different specimens. The effect of electrical discharge is practically absent around 1.8Hz but becomes apparent at lower frequencies. The test was simulated numerically. The simulation captures the effects of electrical discharge but underestimated sensor response amplitude by 40% to 65% for the small and large sensor respectively. High-frequency elastic wave emissions, originating from fibre-reinforced composite failure during monotonic bending, have been measured by an embedded piezoelectric sensor. The signals measured by the embedded piezoelectric sensor have a median signal-to-noise-ratio of 17dB. Using the spectral element method, a parameter study has been performed for a similar, two-dimensional setup. Embedded piezoelectric sensor length and embedment location in the plate thickness has been varied. Both a symmetric and antisymmetric elastic wave have been excited. It has been found that increasing sensor size has a negative effect on the sensitivity of the sensor. The antisymmetric elastic wave is better captured by the sensor response than the symmetric elastic wave. Effects of electromechanical resonance were not noticeable. Monotonic four-point bending failure tests were performed to compare structural integrity of a carbon fibre-reinforced beam specimen with an embedded piezoelectric sensor to a specimen without sensor. With embedded specimens, an average stiffness drop from 3 to 8% is noted. Specimens with small sensors on the tension side of the composite had a decrease lower than 4% in ultimate strain and at maximum 8% reduction in ultimate load and failure toughness. For specimens with embedded sensors on the compression side, the decline was larger, with the larger sensor outperforming the smaller sensor.

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.

Tank containers are intermodal containers for the transport of liquids, gases, and powders as bulk cargo However, hazardous vibration during transportation would generate fatigue damage at discharge area. This thesis aims at calculating fatigue damage at discharge area of a specified tank container.

The calculation is performed in time domain. With FEM models and measured input, hot spot stress is calculated for each hot spot. Probability distributions projects the stress distribution to a longer period thus fatigue damage for 20 years is calculated.

The available data consists of acceleration and velocity of the container during a typical train transportation. A global model and a local model are established with FEM software package. Experimental data from the Lloyd register are used to verify the global model.

In conclusion, during the typical transportation period, fatigue will occur at one of the hotspots.

This master thesis research investigates the possibility to repair carbon fiber thermoplastic aircraft
structures using induction welding. Carbon fiber is conductive and heats up when placed inside an
alternating magnetic field. A generator, coil, and pressure frame are needed to perform induction
welding. The support plates needed to pressurize the carbon fiber parts need to be non-conductive,
non-magnetic, temperature resistant, stiff at high temperatures, and thermally insulating.
The material used is five harness satin weave carbon fiber PPS supplied in pre-consolidated plates
and unconsolidated semi-preg. Three different joint geometries are used in this investigation: a conventional
scarf, a continuous scarf, and a stepped lap joint. The stepped lap joint and the conventional
scarf are milled using a CNC machine. The continuous scarf is produced in a press using specialized
tooling and a press program prescribed by TenCate.
Of each type two specimens are produced: an induction welded specimen and a press joined specimen.
The specimens are tested in tension to determine the tensile strength and stiffness. This gives
the performance of the induction welded joints with respect to the press joined specimens. Additionally,
both of the welded specimens are compared to the pristine specimens.
By measuring the temperature along the weld-line of multiple test specimens, an induction welding
program is obtained for each joint type. Achieving a consistent temperature along the weld-line is
challenging due to the thermal conductivity of carbon fibers and other effects inside the laminate. Similar
to the continuous scarfed specimens, the press joined specimens are created using specialized tooling.
Before testing the specimens in tension, they are scanned using a C-scan. The press bonded
specimens show no flaws, whereas the induction welded specimens do not return the signal to the
transducer. Micrography of the specimens shows similar results as the C-scan. The press joined
specimens show little to no flaws and the induction welded specimens show a significant amount of
voids and small unjoined sections near the tips for the continuous scarf and stepped lap specimen.
The tensile tests show that all induction welded joints perform less than the press joined specimens
with a percentage of about 73-78%. The conventional scarf and the continuous scarf result in a similar
failure strength recovery of about 44% of the pristine failure strength for the press joined specimens.
Both the conventional scarf and the continuous scarf show similar behavior in welding and the tensile
testing, but the continuous scarf is more difficult to produce. Therefore the conventional scarf joint is
preferred over the continuous scarf joint. The stepped lap joint has the best performance, recovering
about 59% of the pristine stiffness. The stiffness of all tested specimens is between 92%-99% of the
pristine stiffness.
Inspection using digital image correlation and finite element modeling indicates the failure initiates
inside the PPS matrix in between the parts at the location of the first or last step. Finite element modeling
also suggests that a smaller angle for the scarf joint or a longer overlap for the stepped lap joint would
increase failure strength.

Thick composite laminates can be used in aerospace and automotive industry replacing metals for weight saving.
Increasing the thickness can bring more defects, like voids.
This master thesis focuses on the multi-scale voids prediction and their mechanical effect simulation of thick composite laminates manufacturing by \ac{RTM}.
It could be potentially used in safe design of thick composite components.
The main objective of the thesis work is to analyze the effect of multi-scale voids caused by the filling process on mechanical properties of thick composite laminates.

First is using PAM-RTM to simulate the locations of voids and their percentage.
Then, the material properties of each \ac{MVE} with different micro and macro voids are evaluated by Digimat at micro and meso scales.
After that, the macro-scale material properties of the thick laminates are simulated in ABAQUS.
%This simulation process is reliable after verifying with the results of a literature.
The simulation processes are verified by using the parameters from the literature and then comparing with the published results.

The simulation from PAM-RTM indicates the average void percentage of the whole laminate reduces with the increasing injection pressure or permeabilities.
The total void volume increases with an increasing laminate thickness but the average void volume does not change with the changing thickness.
The result of this thesis suggests that the effect of multi-scale voids on material properties (i.e., Young's modulus, Poisson's ratio, and Shear modulus) does not change due to an increasing thickness.
The writing presents the actions taken in order to achieve the objective of the research successfully.