The population of orbital debris in Low Earth Orbit (LEO) continues to increase steadily. This situation is driven by a combination of human space activities and collisions between objects in orbit, which are becoming increasingly unavoidable and pose a significant threat to space missions.This work addresses debris–spacecraft collision phenomena by first investigating the hypervelocity impact of a projectile on a single plate. A coupled smoothed particle hydrodynamics (SPH) and finite element method (FEM) numerical framework, implemented in the commercial code LS-DYNA®, is developed to simulate this process and correlated with publicly available hypervelocity impact experimental data. The methodology is subsequently optimised and extended to a more complex configuration, namely a Whipple shield, which is more representative of realistic spacecraft shielding concepts against debris impacts. Validation is performed using an experimental dataset provided by Airbus Defence and Space.Unlike conventional SPH/FEM coupling approaches that are primarily used to improve local damage modelling near the impact zone, the proposed framework is deliberately formulated to enable consistent propagation of shock waves and stress fields into the surrounding finite element domain. This enables the method to be employed not only for accurate fragmentation modelling, but as a physics-driven approach for analysing energy transport and shock propagation within spacecraft structures following hypervelocity impacts.The developed methodology enhances insight into spacecraft structural behaviour under hypervelocity debris impacts and supports its application in the design and optimisation of future spacecraft shielding solutions.

Thermoplastic welded single-lap shear joints made from sandwich-like adherends with a recycled core and virgin face sheets are tested to measure their load-carrying capacity and to quantify damage growth. The results are compared with single-lap shear joints entirely made from laminates consisting of virgin materials. The adherend constituents are T300/polyphenylene sulfide (T300/PPS) and T700/low-melt polyaryletherketone (T700/LM-PAEK), joined via induction and conduction welding, respectively. Test bench data, digital image correlation, a high-speed camera, and microscopy are used to evaluate the structural response of the specimens. Damage initiation and propagation are heavily influenced by the actually welded interface regions. While the virgin joints only experience damage growth along the welded interface, adherend failure can also occur in the recycled core joints. The recycled core joints show a reduced lap shear strength compared to their virgin counterparts, limited to less than 14% for the induction welded T300/PPS specimens and around 40% for the conduction welded T700/LM-PAEK specimens. Despite the reduction in load-carrying capacity, recycled core joints may be considered as a suitable alternative in scenarios where the requirements for mechanical properties are less dominant and environmental considerations are more relevant.

A numerical and experimental investigation is carried out in this study to evaluate the crashworthiness behavior, the energy-absorbing parameters, and the failure mode of thermoplastic composite energy absorbers under axial loading. The energy absorbers are thin-walled circular tubes made of woven polyphenylene sulfide carbon composite and have a bevel trigger on the top. Finite element analysis are conducted to predict the structural behavior of the tubes. To validate the numerical model, four specimens are manufactured and tested under an axial compression load. During the tests, the load and displacement are measured by a load cell, the evolution of the longitudinal strains is captured by digital image correlation, and the progression of the failure is recorded by a high speed camera. The tubes show a progressive failure mode, with delamination between the middle plies. A stable damage propagation is observed throughout the tests, where the bevel trigger plays an important role in increasing the stability and failure progression of the thin-walled tubes. The test results are compared to the numerical prediction, with good agreement for both crashworthiness parameters and delamination behavior. The thermoplastic composite energy absorbers achieved specific energy absorption values of up to 70 kJ/kg, indicating an adequate crashworthiness performance.

Composite interfaces, particularly in joints, play a critical role in the damage resistance and durability of structures for aeronautics applications. This study investigates the use of carbon nanotube (CNT) interleaves for the co-cured joining of composite parts and its effects on fracture toughness and damage progression at the co-cured interface. CNT dispersed in a thermoset resin and partially cured into thin film interleaves at three weight concentrations (0.5% wt., 1% wt., and 2% wt.) of two discrete thicknesses (200 µ and 500 µ) were investigated. The fracture toughness of the co-cured interface with CNT interleaves in mode I and mode II loading conditions was determined through double cantilever beam and end-notched flexure tests, respectively. The results reveal that despite the occurrence of a stick–slip damage progression in mode I, the crack arrest mechanisms and forces are surprisingly predictable based on interleaf thickness. At CNT concentrations above 1% wt., there was no significant enhancement of toughening, and interleaf thickness controlled the crack arrest loads. Damage delay also occurred at the interface due to the activation of multiscale toughening mechanisms. Toughening in mode II was dominated by CNT pullout resistance and, therefore, yielded up to six-fold improvement in critical fracture toughness. These insights offer significant potential for designing joints with nanocomposites for aerospace applications, incorporating inherent toughening and damage delay mechanisms.

Aerospace structures are thin-walled shell structures whose load-bearing capacity is often limited by buckling phenomena. The application of variable angle tow (VAT) composites allows to increase the buckling resistance by tailoring the fiber paths. Fiber placement technologies such as automated fiber placement and continuous tow shearing for VAT composites have been improved enormously in recent years. However, induced material and geometric uncertainties from the manufacturing process have a major influence on the structural performance. The paper focuses on appropriate uncertainty quantification for VAT composites, selecting various uncertainty models based on available data. Different uncertainty models are introduced to quantify the natural variability (aleatory uncertainty) and lack of knowledge (epistemic uncertainty). An uncertain fiber path definition with fuzzy variables is presented to model fiber path deviations. In addition, geometric imperfections are modeled as random fields and as Fourier series to analyze the imperfection sensitivity. Based on this, a design optimization of VAT composites is performed in presence of uncertainties. The introduced methods are demonstrated on a VAT composite panel and a cylindrical shell. Geometric imperfection measurements are provided for the VAT composite cylindrical shell to validate the approach based on experimental results. This paper contributes to a better understanding of uncertainties of tow-steered structures. The results reveal a potential conflict in optimizing the robustness measures (e.g. minimizing the variation of the buckling loads) and enhancing the performance measures (e.g. maximizing the mean value of the buckling loads) visualized by Pareto fronts. This emphasizes the need to consider uncertainties in a design process of VAT composite shells based on multi-objective optimization.

Modern aircraft structures consist of a multi-material mix, dominated by high-performance composites, but also including metal alloys, e.g., for load introduction parts. This experimental research investigates the static and fatigue strength of pinned hybrid titanium-composite single-lap-shear joints. The Ti6Al4V adherend is manufactured by laser powder bed fusion. The joining is done by co-curing with the carbon fiber reinforced polymer adherend. The static tests focus on damage initiation and ultimate load, and are benchmarked by identical joints without pins. The fatigue tests focus on damage initiation and propagation. Digital image correlation is used for damage monitoring. Results show, (i) a high ratio of static ultimate failure to damage initiation load, (ii) early low-cycle damage initiation but then long high-cycle fatigue life until failure, and (iii) the crack stopping effect of the interlocking pins. Furthermore, visual joint failure analysis reveals a variety of damage modes, suggesting comprehensive testing and proper pin design.

To improve the crashworthiness design of composite aircraft structures, analytical models are useful to enable engineers to have a fundamental understanding of the influence of the design variables. As such, during the preliminary design, this knowledge can be exploited, rather than needing to alter an already mature design in a later phase. Accordingly, an analytical model is derived which allows the determination of the mean crushing load and the energy absorption of composite absorbers. The analytical model allows one to accurately predict the mean crushing load of square tube absorbers while altering their side length and thickness. Moreover, the different terms in the analytical model show that the out-of-plane shearing of the material is the major energy dissipating phenomenon. The composite absorbers are then incorporated into a finite element model of the keel section of the thermoplastic composite subfloor of a fuselage demonstrator developed by the Clean Sky 2 STUNNING project. The analytical model facilitates the estimation of the energy absorption and crash load of the fuselage section augmented with the energy absorbers. In this way, during the preliminary design, the absorbers of the fuselage can be designed concurrently for the static loads and for the crash loading, leading to a more efficient design.

Traditionally, launch vehicles are constructed with a series of buckling-prone thin-walled cylindrical and conical shells, in which the buckling behavior of these shells has been well studied and buckling design guidance exists. Conical-cylindrical shell geometry is now being utilized for launch-vehicle stage adapters and payload adapters due to advances in manufacturing and numerical techniques, but there is no available buckling design guidance for this nontraditional combined geometry. In order to provide design recommendations, the buckling behavior and imperfection sensitivity of conical-cylindrical shells and how it differs from the conical and cylindrical components needs to be better understood. From this premise, it is possible to investigate whether or not the buckling knockdown factor guidelines for conical and cylindrical shells outlined in NASA SP-8019 and NASA SP-8007, respectively, are still applicable. The results in this paper will show that the current recommendations are not appropriate in some cases. In addition, it was observed that the large rotations and displacements near the transition between the cone and cylinder can have a larger effect on the buckling load than the presence of radial imperfections for conical-cylindrical shells, which is different than for conical or cylindrical shells. More interesting is the fact that design modifications to increase the buckling capability of a conical-cylindrical shell such as adding reinforcement, which may add mass, will make the shell more sensitive to imperfections. The increased imperfection sensitivity may negate the increase in buckling capability that was thought to be achievable. In the end, it may be more beneficial to design a conical-cylindrical shell in which the buckling behavior is dominated by the more predictable geometric nonlinearity, which may lead to an overall lower buckling load, but a lower knockdown factor may be possible since it will not be as sensitive to the lessknown radial imperfections.

Thermoplastic composite welding is a key technology that can help to make the aviation industry more sustainable, while at the same time enable high-volume production and cost-efficient manufacturing. In this work, characterization, testing and analysis of thermoplastic composite conduction welded joints is performed while accounting for the influence of the manufacturing process. Test specimens are designed from welds of a half a meter long welding tool that is developed to weld the stiffened structures of the next-generation thermoplastic composite fuselage. In the design, special attention is paid to the weldability of the laminates, while ensuring fracture occurs only at the welded interface. Two specimen configurations are evaluated for the Double Cantilever Beam and End-Notched Flexure characterization tests. Moreover, Single Lap-Shear specimens are tested in tension and in three-point-bending. Finally, the characterized material properties are introduced in finite element analyses to demonstrate that the cohesive zone modeling approach can be used to conservatively predict the strength of these welded joints. New insights are obtained in the relation between the manufacturing process, the quality of the weld and the mechanical properties of the joints, which are significantly different compared to autoclave consolidated composites.

In this research, conduction welded C-struts, part of a thermoplastic composite fuselage designed and manufactured in the framework of the Clean Sky 2 STUNNING project, are investigated. Five specimens made of two C-section profiles are manufactured and welded using conduction welding in three different configurations with variations in the direction and distance of the two welded joints. Preliminary numerical analysis using the virtual crack closure technique are conducted to obtain an initial evaluation of the specimens behavior, in preparation of the tests. Experiments are performed under quasi-static loading conditions to measure the strength of the welds. Comparisons with the preliminary numerical analyses show a good agreement in terms of the predicted maximum load, while a clear difference is observed in the initial stiffness, due to the compliance of the support structure. The numerical model is updated, leading to results that closely match the experimental behavior. For all the analyzed specimens, the separation occurs suddenly and no signs of propagation are observed. Experimental and numerical data show no relevant difference in the joint strength among the different conduction welding configurations.

Two curved thermoplastic composite multi-stringer panels with roller boundary conditions are analysed and tested to investigate the buckling and failure behaviour. The panels are made of AS4D/PEKK-FC thermoplastic composite, have five stringers with an angled cap on the side and are joined to the skin with the short-fibre reinforced butt-joint technique. The panels have a roller attached to each loading edge, approximating simply-supported boundary conditions to apply compression and bending. One panel has an initial damage representing a barely visible impact damage in one of the stringer butt-joints, and one panel is in pristine condition. Finite element analyses are performed to predict the structural behaviour, and different approximations of the roller boundary conditions are compared. The analyses include material damage initiation and evolution. The out-of-plane displacement of the panels is measured by digital image correlation, and failure is captured with high-speed cameras. The panels fail in a sudden manner when the cap separates from the web, followed by web failure and skin–stringer separation in the butt-joint. The numerical analysis predicts the overall structural behaviour but cannot capture well the sudden panel collapse due to material damage.

This paper presents the work on six single-stringer specimens manufactured using the card-sliding technique with non-crimp fabrics and adopting a Double-Double (DD) stacking sequence. These specimens, representative of sub-structure level components, are used to investigate post-buckling and failure in aerospace structures. Two specimens maintain a constant thickness cross-section, while four are tapered, two of which incorporate a Teflon insert in the stringer flange. All specimens are tested under compression loading conditions, inducing skin buckling, skin-stringer separation, and eventual collapse. Numerical simulations are validated by experimental results and serve to analyze the specimens behavior and the failure mode. The load versus displacement curves of both experimental tests and Finite Element Method (FEM) analyses are compared, along with the out-of-plane displacement field. Subsequently, the observed failure modes are discussed, focusing on the various mechanisms that occurred and considering the impact of flanges and stiffener tapering. Both the FEM simulations and experimental tests demonstrate good agreement, with the flanges tapering revealing notable results. This offers promising evidence of a viable solution to optimize aeronautical structures and enhance resistance to skin-stringer separation.

This research proposes a computationally efficient methodology using a Constrained Variational Asymptotic Method (C-VAM) for non-linear buckling analysis on a hat-stringer panel with delamination defects. Starting with the geometrically non-linear kinematics, the VAM procedure reduces the three-dimensional (3-D) strain energy functional to an analogous 2-D plate model and evaluates the closed form warping solutions. Utilising the resulting warping solutions and recovery relations for the skin and the stringer, displacement continuity at the three-dimensional level is enforced between the stringer and the skin based on the pristine and delaminated interface regions. Consequently, the constrained matrices obtained from C-VAM is incorporated into an in-house developed non-linear finite element framework. Using the developed formulation, a stiffened panel with delamination of 40 mm between the stringer and the skin is analysed under compression. The results have been validated locally and globally, employing experimental data and 3-D finite element analysis (FEA). Experiments are carried out on the co-cured panel by applying quasi-static loading with displacement-controlled conditions, and 3-D FEA is carried out in Abaqus. Load-response plots have been obtained to validate the results at the global level, and they are in excellent agreement with experiments and 3-D FEA. Subsequently, out-of-plane displacement contour plots are obtained; the number of half waves and wave intensity in 3-D FEA and C-VAM are comparable, although there are minor differences compared to the experimental findings. The proposed framework is shown to be computationally efficient by over 55% as compared to 3-D FEA for performing non-linear buckling analysis on the stiffened composite structure considered in the current work.