The thesis aimed to characterize the effects of manufacturing-induced defects, particularly gaps in variable stiffness laminates. The primary focus centered on designing a specimen capable of accurately characterizing gap effects. A refined, element-level specimen design featuring a controlled gauge zone was developed, based on pre-defined requirements to represent common variable stiffness structures accurately. The specimens were manufactured and tested. The experimental work featured Digital Image Correlation, Dye Penetrant Testing, and Optical Microscopy. Experimental observations revealed deformation peaks induced by gaps, along with trends in void formation and ply deformation during curing processes. Furthermore, digital image correlation (DIC) analysis unveiled distinct out-of-plane tendencies associated with gap locations. Discrepancies between finite element method (FEM) results and experimental strain distributions highlighted the need for improved modeling, particularly accounting for ply waviness. The main hypothesis presumes that the effects of gaps are influencing off-axis plies through matrix failures. The hypothesis was not proven true based on insufficient proof.

The present study, which was carried out in collaboration with GKN Fokker, focuses on incorporating bird strike crashworthiness requirements within a multidisciplinary optimization (MDO) framework. During the preceding three-month internship in the same company, a pivotal contribution to this project was the development of an Abaqus interface for the Multidisciplinary Modeller, MDM, created within the Center of Competence in Design department. MDM is a Python/ParaPy-based automated generator of wings, moveables and flaps, starting from a set of user-specified parameters. The generation of ready-to-run input files thus lays the foundation for the subsequent optimization process, as any changes in materials or geometry can be easily accommodated.

The core objective of the research is to minimize the weight of an aircraft wing while taking into account additional requirements related to the extent of damage caused by bird strikes. Unfortunately, such events occur more frequently than one would be comfortable with, and stringent requirements are set in place to guarantee the safety of the passengers. Among these requirements, the aircraft must be capable of landing safely after such an event, being subject to loads associated with get-home conditions.

As a consequence, two critical constraints are formulated within the optimization framework, addressing the residual strength of the damaged front spar following a bird strike, coupled with a requirement based on a maximum penetration depth. The last constraint has also been included due to the rising popularity of the electric vertical take-off and landing aircraft, which not only fly at low altitudes, thus increasing the risk of bird strike, but may also contain battery packs in the leading edge, for instance, which can pose a significant risk if damaged. To tackle the complexity of this highly-dimensional optimization challenge, a methodology based on Bayesian optimization is proposed, employing surrogate models coupled with a preliminary variable ranking procedure.

The Kriging metamodel is identified as a suitable candidate, thanks to its error prediction capabilities, which are paramount in Bayesian optimization. A variance-based dimensionality reduction method is proposed, which makes use of an initial surrogate to estimate the main and interaction effects of the variables. The quantification of the significance of a variable is expressed as its percentage contribution to the total variance, thus allowing for an intuitive selection of the most important parameters. After the screening procedure is complete, the optimization procedure is carried out in the reduced design space, which uses the constrained expected improvement as an acquisition function. The proposed methodology is then applied on a case study problem, involving a five-bay metallic wing segment subject to the constraints aforementioned, involving 19 design variables representing the thicknesses of various components.

Remarkable weight savings have been achieved, the final result being 40\% lighter than the lightest feasible design among the initial data points. A significant dimensional reduction has also been attained for the maximum depth constraint, which is expected due to the local nature of the impact. Not only did the number of variables greatly decrease from 19 to just 3, but a considerable increase in the accuracy of the corresponding metamodel has also been registered, thanks to an increase in sampling density in the reduced space. However, the variable screening procedure revealed intricate interaction effects with respect to the residual strength of the front spar, emphasizing the nuanced complexity inherent in crashworthiness considerations. Nevertheless, a moderate dimensional reduction has been achieved for this constraint as well, reducing the number of variables to 8, thus proving the efficacy of the proposed variable screening procedure.

In conclusion, the utilization of Kriging models, variable ranking procedures, and Bayesian optimization collectively contributed to the success of achieving remarkable weight savings, proving the efficiency of the proposed methodology. Moreover, it has been shown that the integration of a residual strength requirement is necessary, as many cases were uncovered where no significant penetration occurred, although the application of the considered load case, which is not from critical to an undamaged wing, resulted in high stresses to the front spar of the damaged structure.

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...

Increasing demands for composite structures and sustainability goals have resulted in the growth of thermoplastic polymer matrix composites and automated manufacturing technologies. The in-situ Automated Fiber Placement (AFP) manufacturing of thermoplastic prepreg tapes has the potential to provide a fast and cost-effective manufacturing solution for large composite structures. However, it is still under development because of the complex mechanisms and short processing times involved which leads to several defect formations, especially gaps and overlaps. One of the primary reasons for the formation of these gaps and overlaps is the tape width deformation during placement.

Current literature on tape width deformation shows that the resulting tape width is influenced by several processing parameters such as temperature, pressure and placement speed. However, results from different studies do not agree with each other, indicating that the tape temperature distribution might be at play. Additionally, the conventionally considered tape width deformation mechanism i.e., transverse squeeze flow has been suggested to be incorrect for the AFP process as the experimental deformations do not agree with the results of the transverse squeeze flow model. Therefore, the research objective for this study was to experimentally investigate the width deformation mechanism and the influence of processing parameters for thermoplastic prepreg tapes using in-situ AFP manufacturing and humm3® (from Heraeus) as the heating device.

The specimens were manufactured according to a full-factorial Design of Experiments (DoE) with two settings (high and low) for the following processing parameters: heated length, nip-point temperature and compaction force. The tape width was measured for all the specimens to investigate the influence of the different processing parameters and some post-processing analyses were carried out to understand the tape width deformation mechanism. This included width measurement in the heating phase of the process, surface roughness analysis, tape cross-section profile inspection and fiber-resin content analysis.

From the post-processing analyses and investigations, it was found that the tape width deforms in the heating as well as the consolidation phase of the process. Additionally, the cross-section images show that the conformable roller led the tape edge profile to have a gradual decrease in thickness with a clear slope and the tape edges show a clear indication of spreading of the fiber-resin mixture due to the presence of both fibers and resin. Moreover, the surface roughness data show an indication of the role of temperature distribution because the as-received tape surface roughness was achieved for the higher temperature and longer heated length settings which are assumed to promote better heat distribution in the material.

The influence of the processing parameters on the tape width deformation did not show clear trends for all specimen configurations. However, the exceptions pointed towards the role of temperature distribution in the tape that led to the overshadowing of the effect of other processing parameters. Considering this, it was observed that the change in heated length did not have a significant effect on the tape width except for one configuration i.e., 300 N, 370 °C, wherein a clear increase with no overlap in data was seen. For the effect of temperature and compaction force, it was found that they have an influence on the tape width for temperatures lower than the melting temperature of the polymer resin (Tm). Additionally, the fiber straining effect on the tape width deformation with compaction force was suggested for the higher temperature specimens.

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 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.

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.

Laser-assisted automated fiber placement (LAFP) is a promising additive manufacturing technique for the production of large aerospace components. Thermoplastic prepreg tapes will be placed ply-by-ply on top of a mould and in-situ consolidated. This is beneficial for higher production rates. The mechanical performance of thermoplastic composite laminates is highly dependent on the consolidation quality. One of the quality indicators is the maximum allowable void content after LAFP-manufacturing, which still exceeds the 1% limit for aerospace standards. Challenges remain before LAFP can be completely industrialized with the desired throughput and still obtaining the minimum required quality.

During the heating phase the material is heated to a processing temperature of around 400 ±C within a very short heating time (0.2-0.8s) and no pressure application. As a result of rapid heating, interconnected mechanisms can occur which affect the final consolidation quality. These mechanisms and their relation with processing parameters need to be understood to achieve a high quality laminate manufactured by LAFP. Previous research at Delft University of Technology has shown that deconsolidation phenomena, such as decompaction of the fiber reinforcement network, waviness formation and void thermal growth occur during the heating phase. This will give rise to an increase in void content, surface roughness, out-of-plane deformation and dimensional changes. However, this research did not include tape pre-tension in the experimental setup which is the main component of a LAFP tape placement head. Also, thermoplastic prepreg tapes with a resin-rich surface have been suggested in literature to contribute to a higher consolidation quality of the final laminate. The effect on deconsolidation of thermoplastic prepreg tapes with resin-rich surface during the rapid heating phase is not known yet.

Therefore, the focus of this study is on the effect of a resin-rich surface and tape pre-tension on the deconsolidation response during the heating phase of LAFP. Deconsolidation was quantified through the following response variables (output): surface roughness, maximum out-of-plane deformation, void content, thickness increase and arc-length increase. The results showed that the processing parameters (heating time, heated spot length, resin-richness and tape pre-tension) affect the deconsolidation response through similar interlinked mechanisms as were observed before. However, it has been demonstrated that the mechanisms affecting the deconsolidation response are different due to the presence of a resin-rich surface and as a result of tape pre-tension.

Suprem resin-rich thermoplastic prepreg tapes have a great potential to be used together with the LAFP- process. It has been shown that a resin-rich surface contributes to significantly less decompaction of the fiber reinforcement network. This resulted in less surface roughness and less out-of-plane deformation after the heating phase. Because the fibers are surrounded by resin, no dry fibers are popping-out of the heated surface. The smoother and more resin-rich surface are beneficial for intimate contact development during the consolidation phase of LAFP. It is therefore expected that a higher degree of effective intimate contact can be reached with Suprem resin-rich tapes which is favourable for a higher final quality of a LAFP-manufactured laminate.

Laser heating experiments were performed with three levels of tape pre-tension: 5N, 10N and 15N. Increasing the level of tape pre-tension improved the contact between the tape and the surface of the tool. Since the tool worked as a heat sink in this case, it was shown that local heat absorption occurred only at locations where fiber clusters were present. Increasing the tape pre-tension level towards 10N and 15N seem to be disadvantageous for the LAFP-process. Large out-of-plane deformation was observed and the temperature data showed a highly non-uniform temperature across the simulated nip-point (for 10N and 15N) which is unfavourable for intimate contact development during the consolidation phase of the LAFP-process. It is therefore expected that the optimum pre-tension level lies around 5N. It is assumed that global out-of- plane decompaction and global surface roughness increase will remain lower if a low (5N) pre-tension force is applied.

Over the past decades, fibre-reinforced composite structures have been increasingly adopted into mainstream manufacturing over conventional metals owing to their higher performance attributes, such as superior strength-to-weight ratios. Composite structures offer the unique advantage of tailoring reinforcements in correspondence with design load cases. This allows for more efficient and better performing structures.
Traditionally for composite laminate structures, the design space and freedom were vastly influenced by the accumulated experience in design and certification, as well as available flexibility in manufacturing processes; the available degree of freedom in aligning fibres dictated the design freedom available. With the advent of advanced processes such as automated fibre placement (AFP), reinforced tows can now be placed with more freedom and accuracy; even allowing for tows to be steered during deposition. These are called variable stiffness composite laminates. Fibre directionality within these laminates can be tailored to achieve an optimal load redistribution, thereby increasing its structural performance.
Novel additive manufacturing such as Fused Deposition Modelling (FDM) allow for more complex geometries to be manufacturable in a variety of materials. Such a process can allow more design space and freedom in existing optimization frameworks for designing variable stiffness laminates. Many reinforced thermoplastic materials can be processed using FDM. Short fibre reinforced materials are very readily available and can be used on all commercially available FDM platforms with minimal changes.
This research culminates the three key aspects in engineering – design, material, and process. First, a suitable design framework is chosen, which in combination with added the design freedom by virtue of a novel process such as FDM, is used to design laminates optimized for buckling performance. Additional design freedom is afforded to the optimization framework by means of relaxing the manufacturability constraint – which restricts the maximum allowable curvature of each individual path within the laminate. Secondly, these laminates are manufactured using short fibre reinforced thermoplastic material, for which the shear-induced alignment of the material is analyzed to predict the effective mechanical properties under the parameters used for printing the laminates. Lastly, to validate the effect of additional design space on the effective performance of these laminates, a suitable experimental protocol is devised and used. For the optimized laminates, two cases for maximum allowable steering curvature are considered – one low and one high, and an effective quasi-isotropic laminate is used as a benchmark for comparison. Finally, all the laminates are tested under compression and analyzed. The increase in buckling performance of optimized laminates corresponding to increase in allowable steering is verified, as well as insights are drawn from the processing and experimentation to suggest future recommendations.

The state-of-the-art design approach of composite structures in Formula One (F1) rely on expensive full-scale crash tests and simple models that cannot describe the underlying physical phenomena during crushing. Although the models can describe the crush forces and resulting accelerations, their predictive capabilities are poor, and several iterations are usually required to arrive at a satisfactory design, causing the teams money and resources. Reductions in costs could be possible by reducing the number of full-scale experiments and increasing the amount of analysis.

This thesis aims to evaluate the capabilities of a commercially available material model at two distinct modelling scales by conducting physical and virtual experiments on two different specimen geometries; tubular coupons and a new open-section channel specimen. Firstly, the new open-section specimen with a geometry representing the typical F1 crash structures was designed, manufactured and tested. The newly proposed specimen showed controlled failure modes, and displayed typical characteristics of composite structures under crushing loading. Secondly, the two specimen geometries were numerically analyzed on the macro-scale (laminate-level) and meso-scale (ply-level) using a commercially available material model implemented in the Abaqus finite element software. Mesh-dependency of the macro-scale approach was discussed and numerical aspects required to perform crush simulations on the meso-scale were identified and discussed. Lastly, calibration of the numerical models was performed to understand the effects of model inputs on the output responses, to evaluate the validity of the input parameters, and to assess two different calibration approaches; the first method is based on macro-scale modeling, while the second method is based on surrogate-based inverse parameter identification at the meso-scale.

The numerical models could not reproduce the experimental crush responses of either geometry when using experimentally characterized material data. Following the second calibration approach, the qualitative and quantitative crush responses of the tubular specimen could be accurately described, but the channel specimen showed no improvements in their response. The macro-scale modelling approach with the current material model was judged to be impractical for large-scale analysis, and the meso-scale approach showed need for further developments.

The aerospace industry increasingly uses fibre-reinforced plastics that are manufactured using automated fibre placement (AFP). A common problem in AFP is the occurrence of gaps and overlaps that affect the structural perfromance of laminates. Tow width fluctuation is an important cause of gaps and overlaps. Two developments in manufacturing are combined to create a new mitigation method for gaps and overlaps. One development is fibre steering, which harnesses the flexibility of AFP to vary the fibre orientation within the ply. Another development is smart manufacturing which includes responding real-time to changes. The combination is local fibre steering which compensates for tow width fluctuation by placing tows adjacent to each other. In three steps it is assesed how effective local fibre steering is in increasing the structural performance of composite laminates by reducing gaps and overlaps through compensating for tow width fluctuation in automated fibre placement. Firstly, it is determined to what extent gaps and overlaps are reduced by constructing a conventional and a steered laminate. The tow width is modelled using a cubic spline. The steered laminate uses numerical approximations of the equations for parallel parametric curves. Two problems occurs that are: the accumulation of curvature leading to self-intersections and cusps, and large angle deviations. The curvature accumulation is tackled by using smoothing splines and the fibre angle deviation can be limited by taking a weighted average. Secondly, the tow width model is validated by measuring the tow width of two different materials. Thirdly, the effect of the reduction in gaps and overlaps and the fibre angle deviation on the structural performance is determined by using a finite element method based on the defect layer method that calculated the buckling load and the effective stiffness. Results show that gaps and overlaps cannot be completely eliminated but local fibre steering is able to increase the effective stiffness.

Disassembly of fusion bonded joints aided by induction heating was investigated. Single-lap shear experiments were performed while heating the joint with an induction coil to research the effect on force required and damage inflicted during disassembly. Co-consolidated CF/PEEK samples with and without a metal mesh susceptor and ultrasonically welded samples were tested and compared. An induction heating model was built to facilitate experimental design and to help analyse the effect of the susceptor on the heating process. Induction heating appeared successful in lowering the disassembly force. Large reductions were achieved, but this came at the cost of thermal damage in the disassembled adherends. For a reduction to 37% of the original strength, no thermal damage was inflicted during disassembly, except for one outlier. A susceptor facilitated disassembly for co-consolidated joints, while PEEK energy director remnants hindered heat development in ultrasonically welded joints. Further research is required to develop the method.

Carbon fibre reinforced polymer (CFRP) composites are well-known for their specific strength and stiffness properties. For the past two decades, various industries have increasingly used these materials to improve performance and reduce emissions. This increased use has caused a problematic growth in both production and end-of-life CFRP waste. Recent developments in recycling processes allow recycled carbon fibres (rCF) to have undamaged and clean surfaces with intact sizing, retaining the mechanical properties of non-recycled virgin carbon fibres (vCF) at a considerable cost reduction.
Unfortunately, CF recyclate from these recycling procedures is highly diffuse making them useless for structural application. For making rCF material viable to be used in structural applications, further processing of the recyclate is necessary. To manufacture high-performance rCF materials which feature homogenously distributed, orientated and aligned fibres, considerable effort has to be spent in controlling the microstructure during processing.
The presented thesis has focused on introducing an initial characterisation method that allows for quantitative analysis and comparison between rCF microstructures. Characterisation is performed by analysing elliptical CF footprints in perpendicular cross-section micrographs. Fibres in the micrograph are identified and segmented through a semi-supervised machine learning tool. The elliptical fibre shape is measured by the method of ellipses (MoE) in FIJI ImageJ. This MoE extracts the position and elliptical shape of every fibre footprint in the micrograph. Analysis of the microstructure is done in python to quantify the microstructural morphology by fibre diameter, fibre orientation distribution (FOD), global-local fibre volumes (Vf), fibre density distribution (FDD) and Fibre-to-Fibre (FtF) spacing.
It was found that the method is accurate and reliable in identifying and measuring fibres present in the cross-section micrograph. Similarly, calculated values for Vf, FDD and FtF spacing can be used for characterising the quality of studied materials.
Unfortunately, despite being frequently used in literature to characterise highly aligned and orientated CF materials, all fibre orientation measurements were flawed by the inherent susceptibility of the MoE to calculate fibre orientation with large biases. This orientation bias was caused by the error sensitivity of the arccosines function used for calculating the out-of-plane (OoP) angle. Minute elliptical shape differences in near perpendicular fibres with almost circular footprints resulted in large non-linear errors.
Recommended is to study FODs in samples that are sectioned at a 45 or 60° angle, resulting in large elliptical fibre footprints which are less error sensitive when analysing with the MoE. The induced section angle can easily be transformed back to yield an unbiased FOD measurement.

Due to an ever-growing demand for sustainable transportation, there is an increased need for the development of hydrogen as an energy carrier. Based on mass, hydrogen contains about three times as much energy as conventional gasoline. However, the volumetric energy density, which is the amount of energy per volume unity, is much lower for hydrogen. Therefore, it needs to be stored either at cryogenic temperatures or at hyperbaric pressure. The most common nominal working pressures (NWP) for hydrogen are either 350 or 700 bar. Because of these excessive pressures, the composite pressure vessels (CPVs) require to become thick-walled. This makes the out-of-plane stress components significant which requires calculation models that consider these effects.

This thesis evaluates, whether the use of tow-preg material, could reduce the amount of material necessary to attain a predefined burst pressure. Currently, the production of CPVs is mainly executed using wet winding. The use of this process results in varying material properties due to differences in fibre volume fraction. Also, the friction coefficient is minimal because the resin has a low tackiness. This friction is required for the production of winding paths, which deviate from the traditional geodesic angle. Alternatively, tow-preg material introduces constant material properties and possibly an elevated friction coefficient. This friction coefficient provides an opportunity to vary the winding angle through the thickness, by which the material strength can be made more uniform in every lamina. However, the challenge during the design and production of CPVs is that several effects are present which alter the mechanical response such as varying stiffness on the dome-cylinder interface, non-geodesic winding angle on the dome and resin-rich areas. This thesis will attempt to describe the mechanical behaviour of the CPV, by including several phenomena in numerical models and to minimise the material usage.

Intimate contact development under LAFP-specific thermal and mechanical boundary conditions/interactions and the effect of process parameters are investigated. One-layer, unidirectional strips of CF/PEKK material were placed with different process parameters on a flat tool surface to create different intimate contact conditions. The concept of effective intimate contact, which is based on the resin content at the surface, is introduced and a methodology to measure it from surface micrographs is provided. Degree of effective intimate contact measured from the samples was compared with the existing intimate contact models. The temperature history in the compaction zone was estimated with a finite element model and pressure sensitive films were used to determine the compaction pressure. It is shown that in addition to the squeeze flow mechanism, which is the base for the current intimate contact models, through-thickness percolation flow of the resin needs to be considered to explain the effective intimate contact development.

@en

In this thesis a design optimisation framework was developed, with the aim of reducing the cord usage in TANIQ’s Large Bore Hose product family. To decrease the finite element analysis time, an axisymmetric finite element model was developed using ABAQUS’s rebar method, where the directional cord properties are smeared in hoop direction. The two-dimensional rebar method showed an increased stability and an analysis time reduction of up to 70 % compared to the baseline method. A four-dimensional surrogate model was used in combination with genetic algorithm optimisation to identify the cord angle configuration resulting in the lowest cord usage. The proposed design resulted in a cord usage reduction of 14.9 % compared to a real-life reference design.

Striving to improve structural efficiency the aerospace industry shows increasing interest in variable-stiffness composite laminates. Advanced fiber placement is a hybrid manufacturing technique that offers the flexibility of both filament winding and automated tape laying. With the development of this novel system curved tows can be placed and a spatially variable-stiffness laminate can be designed with continuous changing stiffness from point to point.

The increased design freedom to tailor a structure by in-plane stiffness variation leads to a challenging design optimization problem. A multi-step framework is developed by the aerospace structures and materials department to optimize variable-stiffness laminates. Variable-stiffness laminate design allows for sophisticated designs. Based on this premise it is investigated how such design could improve the structural performance of an engine thrust frame, a structural application that transfers the thrust loads from the rocket engine to the rest of the launch system. The engine thrust frame is subject to cryogenic thermal loads, something not incorporated in the available optimization framework. The goal of this work is to add thermal loads to the laminate analysis routine and to adjust the optimization routine to incorporate thermal influences.

With the thermomechanical optimization framework in place the engine thrust frame is modeled. Conceptual design optimization of the engine thrust frame under thermomechanical loads is performed to increase buckling resistance. A mismatch in the coefficient of thermal expansion is used by the optimal variable-stiffness design. The stiffened areas contract less than the inter-stiffener bay regions. Consequently a stabilizing tensile stress is induced in the prone to buckling bay regions, whereas compressive stresses are distributed to the stiffened areas. Based on the stabilizing thermal stresses and load distribution significant gains in performance are found.

Near free edges of fibre reinforced composite laminates modelled with homogeneous layers, the presence of high gradient interlaminar stresses has been found theoretically. Over a span of nearly 50 years, various methods, including numerical and analytical have been employed for analysis of the high gradient stresses which sometimes exist to an extent of a singularity. The analysis of such free edge stresses is found to be computationally expensive. However, since the interlaminar stresses have been found to play a role in initiating delamination in composite laminates, it becomes imperative to get an estimate of the interlaminar stresses near the free edge to be able to predict delamination initiation through suitable criteria.

The abrupt material discontinuity between dissimilar layers in a homogeneously modelled laminate is believed to be a reason for the appearance of high gradients or singularities theoretically. The mitigation of material discontinuities at such interfaces could be done by modelling the laminate heterogeneously so that the interface is modelled by matrix material and hence, a discontinuity of material could be avoided. This idea of modelling the layers of laminate with explicit definition of fibre and matrix materials has inspired the thesis project with a view to investigate the difference in free edge stresses between the two models and to further draw a correlation between the stresses from the two models to predict delamination initiation.

In previously published literature, the use of average interlaminar stresses up to a certain characteristic distance from the free edge through criterion for delamination initiation prediction is reported. Further it is also found that a single averaging distance can be proposed for a combination of laminates made of the same material. This encourages the idea of determination of averaging distance for a material and find its application on a laminate susceptible to delamination initiation. In the course of the current work, a correlation is proposed between the stresses of homogeneous and heterogeneous model [0/90]s cross-ply laminate of equivalent stiffness made of T300/934 material to determine the averaging distance and apply the determined averaging distance on [±25/90]s laminate to predict delamination initiation due to high gradient free edge stresses.

An averaging distance of 0.125 mm is determined for [±25/90]s laminate made of T300/934 material to find average stresses for incorporation into criterion for delamination initiation. It is found that the determined averaging distance predicts delamination initiation successfully for experimentally reported delamination initiation strain range for the [±25/90]s laminate. Further, the convergence of average of high gradient interlaminar stress is found to be computationally more efficient than the convergence of high gradient interlaminar stress themselves. This indicates that use of average of high gradient interlaminar stresses is an efficient means for predicting free edge stress induced delamination initiation.