Finite Element Methods (FEMs) are widely employed for damage simulation in composites, with the Cohesive Zone Model (CZM) being particularly favored for its robustness and scalability. Despite the CZM's mesh-dependent nature, the Discontinuous Galerkin Cohesive Zone Model (DG/CZM) formulation allows for relatively arbitrary crack propagation thanks to highly refined meshes and its strong suitability for parallel implementation, which is crucial for fracture simulation in composites.

However, a challenge with the CZM in modeling damage in composites lies in determining which damage mechanism occurs at a given numerical interface. To date, fracture simulation in composite laminates has mainly relied on methods designed for isotropic materials, which come with notable modeling constraints such as limiting the arbitrariness of damage evolution and requiring experimental data on fracture propagation.

In this thesis, a novel approach is proposed to ensure arbitrary damage development solely based on the structural stress state using the DG/CZM approach.

Ice accumulation on aircraft surfaces is a prevalent issue, and icephobic coating strategies can be implemented to assist in the removal of or ice on an aircraft. However little is known regarding the effects of novel icephobic coatings and surfaces on ice adhesion strength, which is difficult to measure and quantify, and is typically reported in literature with a high standard deviation and scatter. In this thesis, a reliable set-up for testing ice adhesion strength is successfully designed, constructed, and validated. Using this set-up, it is possible to examine the relationships between various surface parameters and ice adhesion strength, and to analyse the scatter in order correlate the results both qualitatively and quantitatively. Additionally, the influence of material and topology on the adhesion strength and failure mechanisms of ice is investigated.

This thesis investigates the use of Tow-Based Discontinuous Composite (TBDC) interleaves to enhance the mode I fracture toughness of adhesively bonded joints with Carbon Fiber Reinforced Polymer (CFRP) substrates. Aiming to improve joint safety by slowing crack propagation and facilitating less sudden failure, this study focuses on integrating interlaminar-toughened substrates to resist crack growth and enhance the fracture behavior in the joint's substrate. Two main research questions are addressed: the influence of TBDC interleaves on mode I fracture toughness of CFRP laminates and their subsequent effect when used in CFRP-based bonded joints.

For CFRP laminates, Double-Cantilever Beam (DCB) samples were tested across three configurations: a non-toughened baseline and two TBDC-toughened variants. Based on previous research, three DCB configurations identified as the most promising for leveraging TBDC toughening in adhesive joints were tested. The [90/45/-45/TBDC/0]s and [90/60/90/-60/TBDC/0]s laminate substrates were bonded with the low-toughness adhesive Araldite 2015-1, while the [0/TBDC/90_2/0]s substrate was bonded with AF 163-2U, a high-toughness adhesive.

TBDC-toughened CFRP laminates demonstrated up to 130% higher fracture toughness compared to non-toughened counterparts. This was due to TBDC material crack propagation mechanisms such as crack branching, deflection, and fiber bridging.

In adhesively bonded joints, TBDC interleaves in CFRP substrates enhanced the decay of fracture toughness in specimens where cracks deflected from the bond line into the substrate, leading to a less abrupt reduction after reaching peak values. Joints with low-toughness adhesive exhibited more than a 100% increase in crack length from peak fracture toughness to the final value compared to non-TBDC-toughened substrate joints. Meanwhile, joints with high-toughness adhesive demonstrated toughness values 150% to 750% greater than those observed in non-toughened configurations at comparable crack lengths.

These findings highlight the potential of TBDC interleaves to enhance joint toughness, presenting new pathways to improve the safety of composite bonded structures.

To facilitate the transition to liquid hydrogen (LH2) as an alternative aircraft fuel, the stor
age tanks need to be closely monitored to allow for safe operation. This research explores
the feasibility of using an acoustic emission-based structural health monitoring method to
detect and localise simulated damage on a metal double-walled vacuum insulated tank. The
complicated inner and outer tank geometry makes it a novel application of an established
method.

To be able to perform experiments, a metal double-walled vacuum insulated tank with a fixed
and flexible support was cut in half to make the inner tank accessible. Pencil lead break tests
were performed on the inner and outer tank and recorded by lead zirconate titanate (PZT)
sensors to assess localisation accuracy. It was found that a pencil lead break (PLB) performed
on the inner tank could be detected by a sensor placed on the outer tank. Subsequently, an
accuracy of 80% was achieved in making the distinction whether the PLB was located on the
inner or outer tank. For the PLBs placed on the inner tank a 1D localisation accuracy of
17 [mm] and 21 [mm] was achievable for the flexible and fixed support respectively. For the
outer tank a 1D localisation accuracy of 46 [mm] was achievable, while for 2D localisation
this was 41 [mm].

It is concluded that with sensors placed solely on the outer tank, both the inner and outer
tank can be monitored. Additionally, simulated damage on the inner and outer tank could
be localised with an accuracy below 10% of the tank’s dimension. Therefore, based on the
feasibility study performed, acoustic emission-based structural health monitoring is deemed a
viable method to enhance the reliability and safety of metal double-walled vacuum insulated
LH2 tanks.

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

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

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

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

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

The increasing focus on sustainable living and the need to reduce dependency on fossil fuels has led to a growing interest in renewable energy sources. Among these, the wind energy sector has not only experienced significant growth in terms of numbers but also in size. Larger turbines lead to more severe leading-edge erosion and further increase operation and maintenance costs. To mitigate this problem, the new advanced leading-edge protection has become vital in the wind energy sector. This thesis focuses on the evaluation of two polymer coating materials (PA and PD) using a Pulsating Jet Erosion Test (PJET) setup.

The first aim of this thesis is to propose a novel analysis method to address the issue of volume interdependence in the PJET. To achieve this, a concept called "equivalent velocity" is introduced. The equivalent velocity represents the velocity at which a spherical droplet should impact a surface to exert the same kinetic energy per impingement as the actual water slug moving at the impact velocity. By utilizing this concept, the velocity-number of impacts plot takes into account the volume interdependence in erosion experiments.

The second aim is to utilize the PJET to analyze the erosion behavior of PA and PD coatings. The investigation focuses on understanding the relationship between impact velocity and the number of impacts until the incubation period and the breakthrough. The incubation period refers to the interval until the damage is visible and the breakthrough is the moment until the filler underneath the coating is exposed. Additionally, the erosion damage progression of the coatings was analyzed, and the lifetime prediction was evaluated using an existing long-term leading-edge rain erosion model.

The experimental results revealed that the ductile material (PD) exhibits a longer resistance to erosion compared to the stiff material (PA), with the mean number of impacts until breakthrough being 2 to 3 times higher for PD. Moreover, the long-term leading-edge rain erosion model highlights the importance of the accurate measurement of material properties, as lifetime prediction is very sensitive to ultimate tensile strength and Poisson’s ratio.

However, it is crucial to validate the equivalent velocity method through experiments and numerical modeling, while also improving the experimental method to allow for continuous observation of the erosion process in a controlled environment with temperature and humidity regulation. Conducting tests in a wider range of velocities is also recommended. Additionally, improvements for the rain erosion model are necessary to accommodate the utilization of the equivalent velocity.

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.

Efficient maintenance scheduling is a critical objective for aircraft carriers and Maintenance Repair and Overhaul Organizations (MROs) to minimize aircraft downtime. While predictive maintenance models have improved, accurately identifying materials, especially spare parts, for specific maintenance events remains challenging. This paper combines the challenges of identifying spare parts by MROs and aftermarket distributor demand models by developing a robust prediction model (SPSO-CM) to forecast subsequent customer-specific purchases of maintenance planning document (MPD) related spare parts, considering technical documentation and previous sales records. The proposed architecture employs a gradient-boosting algorithm with numerous data mining improvement techniques to predict the likelihood of a subsequent spare part purchase from a customer. A k-means clustering algorithm is used to group spare parts with similar characteristics, as certain specific spare part properties significantly influence demand prediction models. A unique feature selector and nested group K-fold TimeSeriesplit cross-validation method were developed and incorporated into the Bayesian search space to optimize hyperparameters and improve performance. Two test cases were simulated, and the results demonstrated that SPSO-CM is more effective in forecasting proprietary parts and frequently purchased spare parts than those with extremely lumpy demand patterns. Two potential applications for an aftermarket distributor are discussed, one from a customer-level perspective and another at a larger supply-chain level, highlighting its promising capabilities.

During this research, a new approach to predict impact damage was evaluated using the PAL-V Liberty’s propeller blade as the test subject. The purpose of this research was to provide a simple method for impact damage prediction which will support the certification activities of the PAL-V propeller.
This method involves conducting a threat assessment, a Probability of Detection analysis, executing ASTM D7136 impact and quasi-static tests on coupon samples and conducting quasi-static tests on the propeller blade itself. The results from these tests were used to create a two-mass model which can predict the contact force during a propeller impact event. This model was validated using impact test on the propeller blade.
The study conducted a series of ASTM D7136 impact tests and quasi-static tests on coupons, revealing that the coupons exhibited greater load-bearing capacity under impact loading compared to quasi-static loading. These tests also showed that the coupons have an increased stiffness when subjected to impact loading, requiring a multiplication factor of 1.47 to align their stiffness with quasi-static conditions.
The research also involved determining boundary conditions for propeller tests, which led to the development of a custom clamping mechanism. During the quasi-static tests, the differences in stiffness between coupon and propeller tests were identified. These coefficients served as the basis for a predictive
model involving a two-mass, spring and damper system, with parameters derived from coupon tests and the propeller quasi-static tests. After performing the validation impact tests on the propeller blade, it was determined that this model successfully predicted contact forces within 13% of actual test values.
The research demonstrated the potential to predict impact energy requirements for barely visible impact damage, without the use of FEA. The model’s accuracy, while it can be improved, is sufficient to predict impact events effectively, offering a simpler alternative to complex FEA models.

This thesis aims to study the design and optimization of composite wing components for aircraft, focusing on the pivotal goal of reducing weight while preserving and enhancing structural performance. This is one of the goals of the COST Action joined in the span of the thesis: CA18203 - Optimising Design for Inspection (ODIN). Within this COST Action, a comparison between different Finite Element modeling predictions of the response of this structure subjected to 4-point loading has been presented
by Bisagni et al. [1] before this thesis. The thesis, which corresponds to the optimization of this same structure, is the continuity of this work. The foundation of this thesis is built upon a literature study that investigates the optimization of composite wing structures. It explores diverse optimization formulations, including design variables, objectives, and constraints. Different methodologies for optimizing wing structures, like multi-objective and probabilistic methods, are studied. Various algorithms used to tackle these challenges are presented, shedding light on their advantages and drawbacks. Once a solid knowledge of wing structure optimization is built, preliminary analyses (study of coupon geometries, failure criteria, and the buckling behavior of various materials and geometries) form the basis for the upcoming optimization strategy. These preliminary investigations refine our understanding and guide the selection of optimal constraints, such as failure criteria. Based on these preliminary analyses and the literature study, the optimization strategy has been established. A new configuration with cut-outs is developed to explore new innovative designs. To solve the optimization problem, Genetic Algorithms, inspired by real-life behaviors and processes, emerge as the optimization algorithm of choice. These algorithms contribute to the development of design solutions that are better than traditional designs. The objective is to achieve weight reduction in the structure while considering its structural performance. Consequently, the fitness function, employed to assess candidate designs within the Genetic Algorithm, is based on three primary factors: weight, stiffness, and buckling behavior. Weight carries the most significant part of the fitness function, accounting for approximately 80% of the total value, whereas the remaining two factors contribute approximately 10% each. To evaluate these aspects, weight is determined through basic calculations, while stiffness and buckling behavior are assessed by simulating a 4-point bending test using Abaqus.
The implementation of the Genetic Algorithm, adapted to our special case, allows us to showcase its effectiveness in achieving optimal designs. These optimal designs achieve remarkable weight reductions while maintaining great structural performance. Following 20 generations, each comprising 10 potential candidates, one of the two termination criteria is triggered: there is no improvement in the best solution for eight consecutive generations. The optimized solution obtained is 16.89 kg lighter compared to the baseline configuration studied by Bisagni et al. [1]. This represents a significant 25% reduction in structure weight. Concerning the structural performance, only a reduction of 1.5% and 0.25% is observed for the stiffness and buckling performance, respectively, which is acceptable regarding the weight reduction. Nonetheless, achieving such a reduction in wing weight may not be realistic as it was specifically tailored for the test case of this thesis, other loading scenarios need to be studied to assess its applicability.
These outcomes underscore the potential of optimization techniques in aerospace engineering, paving the way for the development of lightweight and high-performing composite wing components.

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

Composite materials have multiple failure modes. Most of these failure modes are dedicated to the different layers, e.g., delamination. Unfortunately, these damages can occur internally, causing it to be invisible to the naked eye. A better understanding of how this happens and what causes it, is required to ensure safety. Therefore, for this research, delamination is the most interesting failure mode. Previous researches have primarily been focusing on quasi-statically loaded UD specimen. However, since a layup in a structural part of an aircraft consists of multiple layers with different fibre orientations, it is not guaranteed that a delamination will occur between two 0º plies. Furthermore, an aircraft is designed to be in service for decades and consequently is loaded thousands of times. A structure which is subjected to a cyclic loading will behave differently than one to a quasi-static loading. Hence, in aviation industries, tests with cyclically loaded specimens are more relevant. This research project focused on fatigue delamination crack growth of multi-directional interfaces. In order to generate mode I delamination, DCB specimens were manufactured and tested. Fatigue experiments were conducted to see the influence of the interface angle as well as that of the orientation of the interface angle. To cover most relevant delamination planes, tests with the following interface were conducted: 0º//0º, 45º//45º, 90º//90º, 0º//45º, 0º//90º, 45º//−45º and 30º//−60º. Furthermore, the crack fronts were monitored by means of C-scans and the fracture surfaces were examined. It was found that for all different interface orientations, fibres bridged. This was demonstrated with an increased delamination resistance for a longer crack. The nature of these bridging fibres, however, differed. Fibre bridging regarding to nesting showed a less significant increase in fracture resistance than bridging fibres due to oscillatory crack propagation behaviour. Furthermore, the different interfaces showed various crack fronts. The crack front tends to propagate towards the direction of the fibres of either of the fibre orientations around the interface.

Wind energy is a growing industry, and in an effort to reduce costs and increase turbine efficiency, rotor blades are becoming increasingly large in size. To facilitate this effort, the SmartBlades2 research project has designed, built, and tested a set of prototype research blades. As part of the SmartBlades2 project, high sensor density modal testing has been conducted on the research blades. The analysis of the modal tests showed good agreement of the global vibration modes with the finite element model predictions. However, the test analysis also identified low frequency vibration modes, referred to as breathing modes, which were not predicted by the finite element models. These vibration modes were found on all of the blades and are characterised by out-of-plane trailing edge panel motion. The objective of this thesis is to identify and predict the aforementioned breathing modes using finite element analysis. To achieve this, three model characteristics are analysed to determine their influence on the breathing mode prediction, namely, model topology, shell element configuration, and material properties. To characterise the affect of model topology, a cut section from the SmartBlades2 prototype blade is modelled with shell elements and continuum element glue joints. To validate the blade section model, a modal test is conducted which identifies breathing modes analogous to the full blade. Various topology features are investigated with the focus on the shell glue joints and spar web joints of the blade section. The analysis shows that while these changes significantly effect the mode shapes and frequencies, none of them predict the experimentally identified breathing modes. To investigate the source of this discrepancy, the modal behaviour of a sample plate structure with the same materials is used to remove the variability of topology. The effects of shell element size and configuration are analysed with mutual comparisons. The analysis shows that higher fidelity element configurations offer no advantage over linear shell elements for prediction of modal behaviour, while the element size shows higher sensitivity. Furthermore, the effects of material properties are examined using the sample plate, subject to modal and flexural tests. It is found that the specified properties are stiffer than measured, and new predictions of the properties are made which better fit the plates experimental results. Finally, the topology, element, and material investigations are then applied to an improved finite element model of the complete blade and correlated with the experimental modal tests. It is found that the improved blade model has closer correlation with the experimental modal tests for global modes, however is unable to predict the identified breathing modes for the blade. It is hypothesised that cause of this may relate to the connection of the spar web with the glue flanges.

To expand the use of additive manufacturing in aerospace towards more critical applications, it is required to design parts in a damage tolerant context. Therefore, the damage tolerance of additive manufactured multiple load path structures is assessed by analysing the fatigue life and damage propagation of components with increasing redundancy. An experimental approach is chosen, whereby specimens with 1, 9 and 81 parallel struts are tested. A decreased fatigue life is found for the specimens with more but thinner struts. This decrease is attributed to manufacturing related effects that occur upon producing smaller elements. The failure of the multiple load path structures showed a step-wise pattern. Due to this, the decreased variation in fatigue life and decreased sensitivity to initial damage, multiple load path structures are more damage tolerant. However, in design a balanced decision should be made upon applying these structures, due to the decreased fatigue life.

Fatigue is the weakening of a material or structure due to cyclic loading and unloading. As such fatigue is very important in engineering fields like aerospace, where structures are exposed to cyclic loads. Therefore theories on the initiation and propagation of fatigue have been extensively researched In the past century. The main Fatigue Crack Growth (FCG) prediction models are based on a Stress Intensity Factor (SIF), ΔK. Representing FCG data as a function of this SIF causes a stress ratio (R-)effect, which is accounted for by means of plasticity induced crack closure. Crack closure causes the crack to close before a zero tensile load is applied, as such influencing the effective SIF. If FCG data is presented as a function of the effective SIF ΔKeff instead of ΔK, the data correlates very well and the R-effect disappears.
However, the theories based on a SIF and crack closure have recently been subject of discussion. The prediction models using ΔK for similitude are empirically derived and thus do not have any physical explanation. Furthermore ΔK has been derived for quasi-static loading conditions and as such it cannot be blindly adopted for fatigue loading conditions. Besides there is a lot of confusion about the phenomenon (plasticity induced) crack closure, load-displacement observations attributed to crack closure could for instance also be attributed to residual compressive stresses. Especially early test results in vacuum environment do not agree with theories based on SIF and crack closure.
An alternative proposed in literature is to approach fatigue from a Strain Energy Release (SER) perspective. This theory approaches fatigue conform the laws of thermodynamics, and claims that not the amount of energy released under quasi-static load conditions should be considered, but the energy released during a complete fatigue load cycle. It also claims that the R-effect is only an artefact of choosing ΔK for similitude. Treating fatigue as a SER dominated phenomenon instead of a SIF dominated phenomenon, might result in a proper description, prediction and understanding of fatigue.
As such the first goal of this research was to investigate if ΔK is a correct similitude parameter for FCG. This was investigated by comparing experimental results in air and vacuum. According to conventional fatigue theories based on SIF, there should be no difference between air and vacuum. Experiments were designed and similar test conditions were applied with the only difference the environment tested in. Experimental results in air showed -as expected- a clear R-effect that could be accounted for by the plasticity induced crack closure corrections proposed in literature. However, the results in vacuum did not show this R-effect, while the results of crack opening experiments and plasticity did not differ compared to results of experiments in air. It was therefore concluded that ΔK is improper to use as similitude parameter for FCG prediction...