The aerospace industry’s push for faster, lighter, and more durable designs highlights the critical challenge of panel flutter in supersonic flight, a poorly understood phenomenon intensified by shock-wave/boundary-layer interactions (SWBLI). This study experimentally investigates the effects of static and dynamic panel deformations on shock-induced panel flutter at Mach 2.0, using open, closed, and ventilated cavity conditions. Results show that cavity closure significantly influences panel response, with the open cavity causing greater oscillations due to strong pressure waves entering it, causing resonance behaviour; closing the cavity seems to filter out this resonance behaviour. Additionally, temperature changes and wind tunnel confinement were identified as factors affecting panel and flow behaviour. The results underscore the importance of understanding experimental conditions to accurately interpret fluid-structure interactions in supersonic conditions.

This thesis investigates the application of hybrid turbulence models for fluid-structure interaction (FSI) to evaluate phenomena such as vortex-induced vibration (VIV) in tube rods within nuclear reactors. The study aimed to identify and validate the most effective hybrid modeling approaches to improve the accuracy and efficiency of FSI simulations in this critical application. Hybrid RANS-LES models aim to use Reynolds Averaged Navier Stokes (RANS) modeling in regions where the flow is relatively steady and well-predicted by the RANS approach, while switching to Large Eddy Simulation (LES) in regions where turbulence is more dynamic and complex. This switch can either be very explicit and segregate the domain into RANS and LES zones or it can be a subtle modification in the transport equations using either RANS or LES as baseline.

The research began with a detailed literature review, leading to the selection of four promising hybrid turbulence models: Improved Delayed Detached Eddy Simulation (IDDES), Scale Adaptive Simulation (SAS), Partially Averaged Navier–Stokes (PANS), and Struct Epsilon (SE). These models were implemented and tested through numerical case setups to validate their performance against established reference data for crossflow over a circular cylinder at a Reynolds number (Re) of 3900.

A comparative analysis was conducted to assess the performance of the selected models. Based on this analysis, two models were shortlisted for further testing. These models were subjected to additional validation on a rigid body motion case to ensure their reliability and accuracy in FSI applications.

In this validation, the models were used to simulate flow over an elastically mounted rigid cylinder in crossflow to evaluate the frequency and displacement of its oscillation. They were tested at different velocities to determine the amplitude and frequency response. The SE model was observed to be more robust and efficient than IDDES.

In conclusion, the research provides validation, selection and comparison of different hybrid turbulence models for FSI applications. Recommendations for future research are also provided to further refine these modeling approaches.

Safety and reliability of aircraft structures are of main concern within the aerospace industry, preferably without influencing the availability of the aircraft and maintenance and operation costs. As damages in composites can be hard to detect, methods are being developed to detect damages which can potentially lead to a (catastrophic) failure of a (sub)system. Interest in structural health monitoring (SHM) has thus gained a strong interest within aerospace engineering, with the development of new systems or optimising existing ones to contribute to a safe and reliable aircraft. This research focuses on the development of an impact damage detection and quantification method for composite structures. The main research objective of this thesis is to investigate the capabilities of a PZT and a state-of-the-art FBG sensor system for passive impact damage detection and quantification on a simple composite panel and a complex stiffened composite structure.

The Flying-V is a novel aircraft concept that has improved fuel efficiency through its distinctive V-shaped structure containing the pressurised cabin, fuel tanks and cargo hold. This configuration presents several key design challenges, one of them being the design of the landing gear. Previous studies indicate that the Flying-V landing gear significantly contributes to the total mass of the aircraft and its placement influences overall aircraft design, both of which impact the viability of the concept. Landing gear designs have been proposed, but are based on outdated requirements, and have notable shortcomings when it comes to feasibility, reliability of mass estimates, aircraft integration and family design commonality. This research aims to explore landing gear designs to identify the best solution for the Flying-V family.

A novel conceptual landing gear design framework is developed, incorporating a gear positioning and sizing optimisation algorithm and physics-based structural sizing. The gear positioning and sizing algorithm allows for efficient exploration of various design features, and can quickly adapt to changing aircraft requirements, making it particularly suitable for the conceptual design phase and the development of family design derivatives. The algorithm is validated using five different reference aircraft with distinct specifications and landing gear configurations, and demonstrated its ability to accurately predict gear positions and lengths. The use of physics-based structural sizing reduces the reliance on statistics-based mass estimation methods, which are known to yield inaccurate results for large and unconventional aircraft. Reference aircraft gear mass data is used to derive a finite-element mass correction equation, improving the accuracy of the mass estimation.

Several Flying-V landing gear concepts are evaluated, each designed with different features, or optimised for a different set of aircraft requirements representing any of the Flying-V family members. For the FV-1000, this resulted in a final design that has a double folding strut, a four-wheel articulated bogie and a strut shortening mechanism, offering a feasible solution with optimal mass and stowage properties. The gear weighs 13.6 tonnes, which is 5.1% of the aircraft maximum take-off mass (MTOM). A derivative design for the FV-900 features the same structural components, but without an articulation mechanism and includes dedicated rolling stock sizing, resulting in a gear weight of 13.3 tonnes, or 5.7% of MTOM. The FV-800 landing gear design follows the same concept as for the FV-1000, but with dedicated structural and rolling stock sizing, significantly reducing mass compared to using the FV-1000-sized common gear. The FV-800 is more constrained by fuel tank capacity and range, which justifies deviating from the commonality principle. The dedicated FV-800 gear weighs 9.2 tonnes which is 5.0% of the MTOM.

Fibre metal laminates (FML) were initially conceived as a hybrid material, aiming to create synergy between the impact resistance of metals and excellent fatigue resistance of fibre reinforced polymers. The purpose of this approach was to overcome the limitations of single-material structures. However, despite its considerable promise, the use of the FML concept has primarily been confined to aerospace applications and heavily relies on synthetic fibres that carry significant environmental implications. Hence, given the growing concerns about climate change and the challenges posed by recycling glass fibre composites, a new generation of FMLs with a reduced carbon footprint should be envisaged.

Research on flax fibre composites reveals convincing mechanical properties and remarkable damping capacities. However, the broader adoption of these composites remains restricted primarily due to issues related to low impact resistance, moisture absorption and flammability concerns. The FML concept presents a viable solution to surmount these constraints, consequently facilitating the integration of these materials into primary structures. Hence, the research endeavour aimed to attain comprehensive insights into FLAx REinforced aluminium (FLARE), particularly focusing on its impact resistance and vibration damping capabilities, which are believed to be the principal benefits of this hybrid material.

The research goal was divided into three distinct research tasks: conducting experimental analyses to characterise the damping behaviour of FLARE, evaluating the impact resistance through experimental means, and validating predictive tools to offer initial insights into the design principles governing such a FML. FLARE, along with flax fibre reinforced epoxy (FFRE) and GLARE specimens, were manufactured using wet layup combined with vacuum bagging techniques.

First, tensile tests were conducted to validate the applicability of the metal volume fraction (MVF) approach in predicting the mechanical properties of FLARE. Intriguingly, the well-known non-linear behaviour exhibited by flax was not observed in the case of FLARE. The results revealed that while the MVF method provided a satisfactory approximation, it was the "inelastic" modulus of FFRE that predominantly contributed to the stiffness of FLARE.

Dynamic mechanical analysis and vibration beam tests were carried out to assess the influence of incorporating metallic layers on the vibration damping characteristics of flax fibre composites. The investigation revealed that the metallic layer predominantly governs the damping behaviour of the FML. Notably, an inverse rule of mixture emerged as the most effective means of approximating its loss factor.

Low-velocity impact tests were conducted to gain insights into the impact response of FLARE in comparison to conventional FMLs. The analysis indicated that the aluminium layers play a significant role in energy absorption, whereas the composite strength emerges as the critical factor influencing impact resistance. A quasi-static analytical model was also assessed, offering an initial estimation of the impact response, yet it warrants further refinement.

In conclusion, the FML concept holds promise for FLARE, but its application requires a novel approach compared to previous methods, to render FLARE viable for practical real-world applications.

Wind turbines experience a wide variety of load cases, both when in operation and when parked. While a turbine in operating conditions often experiences larger loads than in parked conditions, uniquely large inflow angles may occur in parked conditions resulting in less well-understood aeroelastic conditions. Vibrations under these conditions are referred to as stall and vortex-induced vibrations (SVIV). While there has been some success in simulating these conditions, there has been a lack of experimental studies in the literature.

To address this gap, this study analyses the experimental data obtained from a 3.8 MW research wind turbine. This turbine was equipped with strain gauges at the root of 2 blades and in the tower, as well as pressure sensors on one blade and accelerometers in the nacelle and tower. Several experiments were performed with a focus on simulating various conditions where SVIV may occur. This included pitch traverses of a single blade, yaw traverses, and a traverse of the azimuth angle at a yaw angle of 90 degrees to determine the effect of the inclination angle.

Severe stall-induced vibrations were identified at the yaw angle of 110 degrees, when a single blade pointing up in the sky was pitched 180 degrees, and the other two blades were in vane positions at 85 degrees pitch. The severity of these vibrations strongly varied with the wind speed, at an average wind speed of 19.5 m/s the test needed to be stopped early for the safety of the wind turbine, while at 16.6 m/s, the vibrations appeared to reach a limit that was still considered to be safe. The first tower mode experienced the most severe increase in magnitude during these conditions of SIV. In other conditions, no severe case of SIV was identified. However, under several conditions, smaller increases in vibration magnitude were identified. These are likely the result of slightly reduced aerodynamic damping.

No vortex shedding was identified in the data obtained from these experiments. This means that the tested turbine/blade design either did not experience vortex shedding or that it simply could not be measured by the installed instruments. Having pressure sensors installed in multiple locations along the blade would likely help in identifying vortex shedding for future experiments.

Simulations from the experimental conditions were performed using the aeroelastic software PHATAS together with the aero module from ECN (now part of TNO). This simulation used beam-based structural modeling with either a BEM or free vortex wake aerodynamic model. These simulations confirmed the need for realistic turbulent inflow conditions and the need for a good dynamic stall model. However, the structural model failed to simulate all the natural frequencies with the expected accuracy. This should be improved upon for future use of this structural model in parked conditions.

In the last thirty years, massive improvements in flight operations and aircraft technology were implemented, reaching a 54% reduction in specific fuel consumption [1]. However, during the same period, the aviation market has tripled the number of flights per year with a consequent and direct increase in the emissions of greenhouse gasses. To fight against climate change due to the global rise in temperature, international policies are pushing towards a greener future for our society, setting specific objectives for reducing emissions, including aviation ones. In order to achieve the critical goal of climate neutrality by 2050, as stated by the European Green Deal [2], and since the number of total flights is valued will double by that year; significant development steps have to be taken in the aviation industry that has to become greener. The best way to decrease aircraft emissions is to increase its aerodynamic efficiency, so increasing the wingspan of the wing in order to reduce the lift-induced drag. Unfortunately, this has two main problems, the increased total structural weight (that will therefore reduce the payload leading to an increase in the number of flights) and the limited wingspan due to existing airport gate requirements. These issues can be solved using a Hinged Folding Wingtip (HFW) that allows an aerodynamic efficient enormous wingspan during flight operations but will fold on the ground enabling the usage of existing gates. Furthermore, the hinge mechanism improves load alleviation capabilities, particularly during gusts, consequently reducing the bending moment transferred at the root of the wing, allowing beneficial lighter structure. This thesis study focused on a precise aeroelastic characterisation (both aerodynamic and structural) of the hinged folding wingtip device using the instantaneous, non-intrusive and whole-field, large-scale Particle Image Velocimetry (PIV) technique, with the state-of-the-art Shake-The-Box (STB) Lagrangian Particle Tracking (LPT) algorithm in combination with Helium-Filled Soap Bubbles (HFSB) flow tracers and fiducial markers painted on the wing model. This new technique permits the avoidance of intrusive, conventional sensors used for aeroelastic evaluations, like strain gauges, piezoelectric accelerometers, potentiometers, and load cells that alter the structural properties of the tested object but also Pitot tubes that alter the flow behaviour. Furthermore, only pointwise measurements are possible with these devices. Through a wind tunnel campaign at the W-tunnel facility at TU Delft, using a hinged folding wingtip model in both steady and unsteady conditions, it was shown that the implementation of this device shows very promising potential. Not only from an aerodynamic efficiency (due to the increased wingspan) and energy consumption efficiency (due to lower structural weight) points of view but more significantly in the gust alleviation capacities, indeed the peak load and the corresponding root bending moment can be reduced significantly applying a correct timing on the hinge release.

Supersonic panel flutter is a self-excited vibration, where energy is transferred into thin flexible panels, such as skin panels in supersonic vehicles, from the (supersonic) flow around them, which can lead to their rapid fatigue failure. Relevant examples of this include the failure of skin panels in the X-15 vehicle, or violent oscillations in non-adaptable nozzles during start-up, such as the space shuttle main booster. With the contemporary trends of decreasing weight and increasing operational speeds in the design of next generation aircraft and launch vehicles, such structures will become more common, and flutter needs to be inherently accounted for in the design.

In recent years, advancements in numerical techniques, such as Computational Fluid Dynamics (CFD) and the Finite Element Method (FEM), have led to numerous studies that address the effect of non-linearities, such as flutter behaviour in the transonic and hypersonic regime. In spite of that, experimental data needed to validate these studies is limited, based on coarse pointwise measurement techniques, and only focused on the structural response of the panel rather than on the entire Fluid Structure Interaction (FSI). The use of simultaneous, full-field and non-intrusive measurements can be used to address these issues. To this extent, panel designs were developed and an experimental campaign was proposed where Digital Image Correlation (DIC) and Schlieren are simultaneously used by means of high speed cameras to study classical panel flutter at Mach 2 inside the ST-15 wind tunnel at the Delft University of Technology. Additionally, a Laser Doppler Vibrometer (LDV) was used to validate the DIC measurements. The experiment allowed for a comparison with well understood theory, such that the setup can be proven to be robust, and can therefore later be used to assess the effect of panel flutter non-linearities or other supersonic FSIs.

Synchronisation between DIC and schlieren was demonstrated through a matching frequency spectrum, and maximum correlation inside corresponding pixel windows between panel displacement and schlieren grey-scale fluctuations with a zero time-lag. One finding was that during the panel down-stroke phase, a travelling wave character was observed which has previously been found only under transonic conditions. Furthermore, at moderate dynamic flutter pressures, a typical second bending panel flutter mode shape was observed, which grew into a single bending mode again with increasing higher dynamic flutter pressures. From the spectral analysis it was found that all configurations fluttered at the same frequency of 770 [Hz], independent of dynamic flutter pressure or edge conditions. This was higher than expected from analysis. These combined observations led to the conclusion that, in the current facility, a frequency lock-in occurs due to merging of the flutter phenomenon with a wind tunnel resonance vibration.

These observations provide an insight into combining these experimental techniques to study supersonic FSI behaviour, and particularly into certain unexpected behaviour encountered at the TU Delft ST-15 wind tunnel. This can be used to further tune future experiments and guide researchers to obtain a better understanding in the non-linearities found in panel flutter.

Several structural models that could serve as building blocks for a next-generation FSI-based kite design tool have been developed. The models represent the leading-edge inflatable V3 soft wing membrane kite of Kitepower B.V., by using a wireframe multi-plate representation. Where the shape deformations are calculated using bridle line system models. Both by using a trilateration algorithm and a particle system model were accurate results found. Accurate in comparison to the experimentally obtained data from a photogrammetry analysis. The existence of slack and its non-dependence on empirical relations made the force-based particle system model best.

As of April 2019, upset prevention and recovery training in flight simulation training devices is a mandatory practice for commercial and civil aircraft pilots. Aircraft stalls are a well-known upset type, therefore simulation of aircraft stall behavior is required. A key characteristic of stalls is the buffeting component, which in current stall models is still insufficiently modeled. In this research, a new methodology to more accurately model stall buffet behavior using swept
wing flight test data is presented. Buffet effects occur after exceeding the critical angle of attack, with an aircraft type-specific buffet onset duration to fully develop maximum buffet intensity. The buffet transient behavior is modeled with a frequency response fit and a multivariate second-order polynomial to capture aircraft eigenmode-shape frequencies and buffet intensity respectively. Aircraft recovery and thus receding buffet effects occur at an increasing angle
of attack, which is used as buffet offset. Generalization of the results was shown with the validation of data set of a straight wing aircraft, which indicates a step towards a more generic stall buffet model methodology.

During aircraft design,multiple tools are utilised to inspect the performance of the configuration. As the design matures, higher fidelity analyses are conducted to predict the flight dynamics of the aircraft. These analyses are conducted by using semi-empirical relations, numerically analysing flow behaviour, conducting wind-tunnel tests and performing scaled test flights. However, Semi-empirical relations might not hold for next generation aircraft and wind tunnel testing and scaled test flights are extensive and are also prone to accuracy issues. A best of both worlds can be found in numerical analysis. However, an increase in flow fidelity modelling comes with an increase in computational cost. Besides, complete analysis of all possible manoeuvres of a design increases the number of computations significantly. Current methods cope with this issue by using flight dynamics models based on so called stability derivatives, instantaneous values which couple flight state parameters to aerodynamic loads to predict aircraft flight dynamics. However, these models do not take into account time dependency. Therefore, these methods do not accurately predict the flight dynamics of agile aircraft, such as unmanned combat aerial vehicles, undergoing rapid manoeuvres where unsteadiness dominates flow behaviour. This conventional reduced-order modelling method, in which samples of the full-order model are taken in the form of stability derivatives, causes design iterations to be analysed inaccurately. The objective of this report is to investigate reduced-order modelling for flight dynamics prediction, thereby comparing conventional techniques to a method which does take into account unsteadiness in flow behaviour.
The method investigated is based on indicial step response functions, which are samples in the form of unsteady aerodynamic flow behaviour functions of the full-order model. The idea is that once these samples are known, any flight manoeuvre can be analysed within minutes. Research found in literature has assessed some of the capabilities and limitations of this method, but not yet applied this to flight dynamics prediction. The research described within this report will address this gap by using two test cases. The first testcase is used to assess the assumptions made in literature, on aerodynamics loads modelling, by applying the method on a two-dimensional airfoil in subsonic flow conditions. It was found that the indicial step response functions are indeed representing the full-order model, thereby taking into account unsteady flow behaviour in aerodynamic loads prediction. In longitudinal motions, the angle of attack and pitch rate effect need to be taken into account to predict lift, drag and pitching moments. Multiple frequencies of the same manoeuvre can be analysed within minutes once the samples are calculated. Results show that the accuracy of the predictions becomes a trade-off issue between samples calculated and accuracy required. The second testcase is used to apply the indicial response functions to flight dynamics prediction of an agile
unmanned bomber aircraft undergoing fast manoeuvres. A longitudinal-directional climbing manoeuvre was calculated by developing a flight dynamics model based on stability derivatives. The flow behaviour encountered during this manoeuvre was analysed to include highly unsteady and non-linear phenomena (e.g. vortices and flow separation) at higher angles of attack. By comparing the results of themethod under investigation to the full-order solutions, it was shown that aerodynamic flight dynamics predictions were accurate in capturing unsteady behaviour and weak non-linear flow behaviour. However, the samples proved to be inaccurate in representing behaviour in highly non-linear regions. Concluding, this means that indicial step response functions provide more accurate flight dynamics predictions than conventional stability derivatives in representing unsteady flow behaviour. The accuracy of the predictions are highly dependent on the samples chosen. Several samples suffice to predict the unsteady behaviour for linear and weak non-linear flow regions of the flightmanoeuvre. If surrogate modelling is applied, the method can become more computational efficient than conducting multiple full-order time-marching numerical calculations. It is recommended that more research is performed on indicial step response functions
in capturing highly non-linear flow behaviour, as the research showed that the size of the samples affects the flow behaviour representation.