As the need to reduce the carbon footprint of the energy sector grows, wind energy has emerged as a leading renewable source for sustainable power generation. In particular, the offshore sector has gained significant attention due to its higher wind speeds as well as its reduced noise impact. TU Delft's Faculty of Aerospace Engineering has identified accelerating the energy transition towards renewable energy sources as one of its main goals. To achieve this, offshore wind turbines must be tested under realistic Atmospheric Boundary Layer (ABL) conditions.

Wind-tunnel ABL simulations typically use spires, barriers and roughness elements as passive devices. However, this configuration requires a long downstream length for the boundary layer to develop. Therefore, an alternative approach developed by Cowdrey (1967) is used, which employs a non-uniform rod grid to produce a prescribed power-law mean velocity profile. Moreover, unlike spire-based configurations, this method allows for variations in the turbulence levels of the resulting boundary layer through the grid design parameters.

This thesis investigated the capability of Cowdrey's method to reproduce an offshore ABL in a short open-jet test section. Particle Image Velocimetry and Hot-Wire Anemometry were used to measure the mean velocity profile, turbulence intensity, integral length scales and power spectral density. Cowdrey grids were evaluated over a test matrix with four rod diameters, two cross-sectional shapes, two turbulence design settings, two target boundary layer thicknesses at three free-stream velocities.

Within the tested velocity range, the use of Cowdrey grids increased the boundary layer height by up to 26 times compared with the naturally developing layer within the same fetch length. The fitted power-law exponents for most configurations were slightly larger than the design value but remained close to the target value overall. Moreover, most grids exhibited a good agreement with the logarithmic law in the inner layer and within the available downstream distance, attained the characteristic aerodynamic roughness length of an offshore Atmospheric Boundary Layer. The measured turbulence intensities exhibited elevated near-wall levels relative to semi-empirical relations but qualitatively followed the predicted offshore ABL trends. The power spectral density conformed to the von Kármán model across all sampled heights, with noticeable departures only near the top of the outer layer at the furthest downstream station. Lastly, the longitudinal integral length scales tended to be consistently lower than predicted by semi-empirical relations.

The study provided valuable insights into the applicability of Cowdrey’s method in reproducing a scaled offshore ABL. It laid the foundations for further research regarding ABL testing in TU Delft's experimental facilities. Further improvements should focus on bringing the fitted power law closer to the target value and refining the turbulence characteristics to minimize deviations from accepted standards. ... Read more

Acoustic liners have been extensively studied from a purely acoustic perspective. Recently, the interest in characterising their aerodynamic performance has increased, as they are known to provide an increase in drag, compared to their respective smooth counterpart, up to 70%. The goal of this work is to experimentally test various acoustic liner samples to study this drag contribution. The hypothesis is that it is possible to decrease the added drag without having to compromise the acoustic performance.
A variety of orifice cross-section configurations are investigated. Samples are tested in different experimental campaigns, combining the use of two different wind tunnels (the Delft University- Boundary
Layer Facility in Delft and the Flow Duct Facility at NLR, Marknesse) and two different experimental measurement setups, both to investigate the setup effects and to provide a deep understanding of liner performances in different operating conditions. All sample porous plates are 3D-printed by means of Stereolitography (SLA). Direct drag measurements are performed to obtain the aerodynamic response. Insertion Loss measurements are carried out to compare the acoustic attenuation of the different samples.
The direct drag measurements show an increase in added drag for all samples in all tested conditions,
with 0% < ∆CD < 37%. Compared to the conventional cylindrical orifice shape, two configurations show a lower added drag: the tapered and the inverse chamfered. Due to shape simplicity, which makes it easier to manufacture with different techniques, the tapered is taken as reference ’improved’
configuration. The relative drag reduction caused by the tapered configuration ranges from 20% to 40%, with consistent results between different campaigns. The insertion loss measurements, on the other hand, show no significant difference between different samples. All samples have a low acoustic
attenuation in the no-flow case, reaching a maximum of 2.5dB. With increasing M, the attenuation increases in the whole frequency range, probably due to the testing configuration and wind tunnel effects.
The measurements show important novel insights on the drag increase caused by acoustic liners. Data from different campaigns, facilities and measurement systems compare positively and a possible model is proposed, which relates the drag increase to the product of the Open Area Ratio σ and the inner-scaled pore diameter at the top of the liner porous facesheet, d+. It allows to estimate the drag increase,
given the geometric parameters of the liner and the flow conditions (in terms of a viscous parameter, e.g. uτ). This model does not take into account the effect of the material surface roughness, which has
a considerable impact, due to the contribution of the ’pure skin-friction drag’ on the total drag. From the acoustic measurements, no significant difference is found between the acoustic attenuation of all
liner samples, showing that a lower drag increase can be achieved without sacrificing the acoustic performance. All in all, further development in acoustic liners is considered crucial to achieve optimal noise attenuation as well as aerodynamic performance.
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This thesis presents the first systematic experimental study of boattail radius effects on Hammerhead Fairings (HHFs) in transonic flow, using the Coe and Nute Model 11 as a representative platform. Boattail radii were selected based on ratios of radius to boundary-layer thickness, and tested in two campaigns: high-speed schlieren imaging, oil-flow visualisation, and PIV at selected cases. Results show that increased radius generates a third expansion region, often with a lambda-shock, reducing separation length by 10–25% but increasing reattachment-point oscillations. Strong radius effects promoted shock-induced separation and lifted the shear layer at the shock foot. Two dominant unsteady modes were identified: a flapping mode linked to reattachment motion and shock features, and an undulation mode related to momentum injection and ejection from the separation bubble. Findings hint at trade-offs between separation reduction and increased unsteadiness, with implications for launch vehicle stability and structural loading. ... Read more

Reducing aerodynamic drag is crucial for lowering energy consumption and emissions in industries such asaviation and fluid transport through pipelines. In cruise conditions, skin friction drag is the dominant contributor to aerodynamic resistance and has therefore been the focus of numerous studies. A widely discussed technique for skin friction reduction is active forcing through spatio-temporal spanwise wall oscillations, which has been numerically shown to achieve drag reductions of up to 50%. This technique relies on creating a spanwise wall motion, generating a phase-varying spanwise velocity profile that suppresses near-wall turbulence. However, practical implementation remains challenging due to mechanical complexity and high energy costs, leading to interest in passive alternatives. Among these, dimples and Oblique Wavy Walls (OWW) are examples, as their surface deformations have shown to introduce a spatially varying spanwise shear layer into the flow. Through geometrical surface deformations pressure gradients are formed, which give rise to a spanwise component of the flow in the near-wall region. Although these techniques have demonstrated meandering flow structures, no significant drag reductions have been established previously. This motivates the investigation of sinusoidal undulations as a potential passive method for introducing a spatially periodic spanwise velocity component.

This thesis studies the flow topology of sinusoidal undulations to assess their ability to impose a spanwise shear layer. The geometry is characterised by a streamwise oriented sinusoidal groove, that displaces in the spanwise direction, having a dimple-like cross section. The influence of the geometry is assessed in terms of, spanwise displacement amplitude A, depth-to-diameter ratio d/D, and scale (while retaining geometric similarity). The flow over sinusoidal undulations was visualised by means of Particle Image Velocimetry (PIV) in (i) a stereographic PIV in the streamwise-wall-normal plane to quantify the spanwise velocity amplitude and the penetration depth, and (ii) a planar PIV in the streamwise-spanwise plane in the near-wall region (y = 1 mm) to examine the near-wall flow structure. These experiments were conducted in the M-tunnel of the Delft University of Technology.

Analysis of the experiments confirmed an alternating spanwise movement of the flow that aligns with the geometrical spanwise angle of the sinusoidal undulation. Simultaneously, the windward edges of the geometry act as diffusers, pushing part of the flow over the edge and out of the groove onto the flat area. These combined effects led to the identification of a converging-diverging flow structure, locally resembling dimple flow structure. Additionally, a relation was found between the undulation geometry and flow characteristics such as maximum spanwise velocity and penetration depth. While spanwise velocity oscillations were observed across all geometries, a larger geometrical displacement intensified the diffuser effect of the windward edges. The measured velocity magnitudes were comparable to other passive techniques, but lower than those achieved by active spatially periodic forcing with optimal actuation conditions for drag reduction. Furthermore, these velocities were associated with penetration depths into the buffer layer.

This study demonstrates that sinusoidal undulations are able to introduce a phase-varying spanwise velocity of w⁺≈ 1, or approximately 4% of the freestream velocity, forming a converging-diverging flow topology, bearing resemblance to dimples with comparable spanwise velocity amplitudes (w/U_∞ ∼ 3%). The measured spanwise velocity profiles exhibit similarities to the analytical solution of the spatial Stokes layer, though the penetration depth follows a different definition. The near-wall shear layer in the viscous layer resembles a horizontally mirrored Stokes layer, and is characterised by a secondary deeper penetration resultant from the imposed pressure gradient. To further assess the drag-reduction potential of sinusoidal undulations, drag balance measurements and numerical simulations are recommended. ... Read more

An impinging Shock Wave-Turbulent Boundary Layer Interaction (SWTBLI) at Mach 2 was investigated experimentally while implementing two-dimensional Shock Control Bumps (SCBs). The aim was to investigate the changes in the unsteady dynamics while changing the bump ramp angle, tail angle, spanwise shape, and the impinging shock location. Schlieren and oil flow visualisations were used to identify these changes. Unsteadiness was quantified through a spectral analysis based on Welch’s method, and a Canny-based edge-detection algorithm was developed to track the impingement and separation location in the Schlieren images.

An uncontrolled SWTBLI could successfully be generated, and the implementation of the baseline bump showed a replacement of the unsteady separation shock by a steady compression ramp shock originating from the leading edge. Spectral analysis confirmed this behaviour since the characteristic low-frequencies of the uncontrolled separation shock seemed to be removed for this compression ramp shock. An increase in the bump ramp angle showed the progressive generation of a separation shock upstream of the bump with spectral content trending towards the low frequencies of the uncontrolled interaction. On the contrary, no alterations were observed for an increase in tail angle. An upstream impingement produced similar behaviour as the increase in ramp angle; a separation shock was generated upstream of the bump with increased low-frequency spectral content. The downstream impingement, however, did not show any alteration in the same region. Major factors influencing the unsteady dynamics are identified as the impingement location and the ramp shape of the bump.

The developed edge-detection algorithm proved unsuccessful in quantifying the unsteadiness although it could detect the impinging and reflection shock of the interactions. Spatial standard deviation distributions of the interaction revealed increased deviation values in the impinging shock suggesting that the impinging shock fluctuates. Rather, it is suspected to be a form of noise intrinsic to the experimental technique. Therefore, this affects the detection of the edges and the calculated impingement and separation location. Future research is suggested to improve the noise mitigation method in the algorithm. Additionally, the benefits associated with the 2D-SCB would be most noticeable in an integrated approach with an industrial application. Finally, numerical simulations and/or different experimental techniques are suggested for future research to obtain a better quantification of the interaction and unsteady dynamics.
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The 'hammerhead fairing' (HHF) launcher configuration uses a payload fairing that has a larger diameter than the launch vehicle. This allows for larger payloads to be carried using the same launcher components. However, the HHF leads to high aerodynamic forcing in the transonic regime called buffeting. If unaccounted for, buffeting can lead to launcher failure. The flow phenomena associated with buffeting were investigated on the Vega-E launcher, which is currently being developed by the ESA, using Pressure Sensitive Paint (PSP) in the TST-27 tunnel.

PSPs are a relatively new flow measurement technique that allows for quantitative measurements of surface pressure levels in a non-intrusive manner. High spatial resolution, which is equivalent to using millions of pressure taps, can be achieved using PSP. This allows for CFD-like surface pressure resolution to be obtained from wind tunnel experiments. However, PSPs had previously never been successfully implemented in Delft. Therefore, a test set up was devised for experiments in the TST-27. The final implementation was found to have a pressure uncertainty of at most 2.5 kPa, with a spatial resolution of 0.088mm.

First, validation measurements were done on the Coe and Nute Model 11. The Model 11 is a common HHF test case in literature that is used to study the buffeting phenomenon. Results obtained in the TST-27 were compared to historical data and showed good matching when accounting for blockage. After the validation, measurements were taken on the Vega-E launcher at different Mach numbers and angles of attack to quantify the effects of each variable on the pressure field. Lastly, the impact of protuberances on the model was studied. The protuberances that were used mimic those that will be present in the full-scale launcher. ... Read more

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.

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This thesis investigates the role of winglet slots in reducing tip vortices, inspired by bird wingtips. The effects of slotting and wingtip flexibility were analyzed using multiple bio-inspired models tested in a low-speed wind tunnel at Reynolds numbers from 2.9 × 10⁴ to 8.8 × 10⁴. Three experimental techniques were employed: force balance for aerodynamic force estimation, motion tracking of wingtip movement, and stereo PIV analysis to measure cross-flow planes in the winglet wake. Non-intrusive methods were used to estimate drag components based on PIV data and wake integral approaches. The slots significantly reduced induced drag by breaking down the tip vortex into smaller, weaker vortices. This breakdown was visualized using s-PIV results and analyzed through velocity component contours. The streamwise vorticity breakdown was examined in relation to wingtip flexibility, and the vortex cores' location and characteristics were analyzed, along with the phenomenon of vortex wandering.

By the end of the thesis, it was concluded that although slotted winglets reduced induced drag, they led to a significant increase in profile drag, resulting in higher overall drag compared to unslotted winglets. Additionally, the slots caused a loss in lift. The overall aerodynamic performance, measured by the lift-to-drag ratio, was also higher for the
unslotted winglet. Among the slotted winglets, static bending was common, while wingtip vibrations were minimal. The most flexible slotted winglet had the lowest induced drag. However, the results indicate that at very low Reynolds numbers, slotted winglets deteriorate the wing’s aerodynamic performance compared to unslotted ones. Among the slotted winglets, those with intermediate flexibility demonstrated the best aerodynamic performance, highlighting the importance of optimizing wingtip flexibility. ... Read more

An important challenge in the electrification of aircraft propulsion systems is the design of thermal management systems because of an increased heat load that needs to be dissipated. As an alternative to high-drag external heat exchangers, one can consider surface heat exchangers to dissipate the extra thermal energy. This reduces the size of the external heat exchangers and consequently reduces drag. However, a non-adiabatic wall can affect skin friction through a movement of boundary layer transition. Limited studies are available on the effect of non-adiabatic surfaces on laminar-to-turbulent transition in swept wing boundary layers dominated by crossflow instability (CFI). Therefore, the current work experimentally investigates the effect of surface heating on the stability and breakdown of the stationary crossflow instability. The experimental work is supported by Compressible Linear Stability Theory (CLST) computations.

Hot-Wire Anemometry (HWA) and Cold-Wire Anemometry (CWA) measurements of the boundary layer are performed on the STEP model, which features a 45 degree swept flat plate, for both adiabatic and heated surface conditions. Both the experimental and CLST results show a destabilisation of the primary instability linked to the increase in the growth rate of the stationary crossflow (CF) mode. The experimental results show that the type-I secondary CF instability exhibits a larger magnitude in the presence of wall heating and the mode emerges upstream compared to the adiabatic wall condition. The type-III mode displays a significant increase in magnitude in the presence of wall heating, thereby indicating a considerable destabilisation. The effect on laminar breakdown is identified by analysing velocity fluctuations in the 12-17 kHz frequency band in planes parallel to the surface for two different wall distances. A temperature ratio of 1.035 is found to advance breakdown by 5.7%. ... Read more

This thesis, divided into two parts, explores shock interaction phenomena occurring in high-speed platforms. part one focuses on shock-shock interactions (ssis), which manifest as a bi-stable system with either the regular interaction (ri) or the Mach interaction (mi) as outcomes. Through numerical simulations, both interaction types are examined in the presence of perturbations to provide insights into their stability characteristics and the riÕmi transition process, which exhibits hysteresis effects. Subsequently, attempts are made to replicate these hysteresis effects in the transonic–supersonic blow-down wind tunnel (tst-27) at tu Delft by continuously varying the free-stream Mach number during a run. The evolving shock system is continually tracked using a systematic post-processing methodology that integrates schlieren visualizations, synchronous pressure readings, and insights gained from a variable focal plane study conducted with a focusing schlieren system... ... Read more

Wind tunnel investigations of cycling aerodynamics are conducted in air streams that are designed to have minimum turbulence. This is not representative of the true, turbulent wind environment encountered by a track cyclist in an indoor velodrome. The goal of this thesis, broadly speaking, is to assess the effects of this turbulence on cycling aerodynamics.
To address this question, two experimental campaigns were conducted. The first consisted of track measurements, to measure the turbulence intensities and spectra encountered in an indoor velodrome, by an isolated cyclist and by a cyclist in the far wake of another. This turbulence data was collected using a three-hole pressure probe. In the case of an isolated cyclist, it was observed that the airflow was dependent on the cyclist's location on track. A spectral analysis also revealed the presence of periodic components to the airflow, a result of the cyclist's cadence. When riding in the far wake of another cyclist, the turbulence encountered was stronger when the distance of separation got smaller.
The second experimental campaign was conducted in the wind tunnel. First, the measured track turbulence intensities and spectra were simulated using grids, for small scale, high frequency turbulence, and large obstacles, for large scale, low frequency turbulence. Once the simulated turbulence matched the track turbulence to acceptable levels, its effects on the drag of a cylinder model, representative of the cyclist's limb, were assessed. These drag measurements were conducted using an external force balance. It was found that for a smooth surfaced cylinder, free-stream turbulence was enough to trigger flow transitions and cause a reduction in the measured drag. For a rough surfaced cylinder, increasing the turbulence triggered flow transition earlier, and led to an increased Reynolds number range for the critical flow regime. It was also seen that large scale turbulence did not have a significant effect of the drag of the cylinder.
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TOP contoured nozzles, with large area-ratios, are commonly employed in rocket propulsion systems as they feature an excellent thrust-to-weight ratio. A significant shortcoming to this design is that, during the startup and shutdown transients of a LRE, the internal nozzle flow progresses through a series of overexpanded flow states - Free Shock Separation (FSS) and Restricted Shock Separation (RSS), which produce critical loads associated with SWBLI and asymmetric flow separation. Exacerbated by FSI, this operational phase is known to generate the highest vibroacoustic loads at which payload and vehicle structures are subjected to when the engine is operated at off-design conditions.

Motivated by the importance of understanding how the interaction between the developing flow and the vibrating nozzle walls has an effect on supersonic noise generation and propagation, this work has studied the effect that wall compliancy has on the vibroacoustic loading of TOP contoured nozzles. This is demonstrated by means of cold flow tests carried out in the High Speed Laboratories of the Delft University of Technology on a stiff-walled aluminum nozzle, which serves as a baseline test case, and on a urethane-based compliant walled nozzle. Tests are conducted under comparable flow conditions and test parameters are measured by means of acoustic and optical techniques. Simultaneous recordings are performed and include the nozzle-wall deformation, by means of stereoscopic tracking of tracers on the nozzle lip, the imprint of the near-field acoustic signature, by means of arrays of pressure-microphones, and Schlieren imaging of the jet plume.

Measurement data allows for a Fourier decomposition of the nozzle lip displacement and of the acoustic pressure field in azimuth. Reconstruction of the instantaneous plume development enables the identification of the main flow structures responsible for noise generation.

Comparison of results between the two test articles highlights a different spectral content and directivity pattern. Correlation between the structural displacements and the acoustic signal, together with the use of DMD, quantitatively aids the investigation of how FSI has an impact on the generation of an aeroelastic tone at 180 Hz. Findings suggest that its production is the result of the periodic thickening and thinning of the shear layer owing to the heightened flapping motion of the nozzle lip preceding RSS transition, driven by an intensified shock foot instability. ... Read more

Turbulence is a ubiquitous phenomenon in nature, characterized by complex and unpredictable fluid motion that is challenging to simulate. Turbulence modeling seeks to develop mathematical frameworks capable of predicting turbulent flow behavior in various systems. Common approaches include Reynolds-Averaged Navier-Stokes (RANS), Large Eddy Simulation (LES), and Direct Numerical Simulation (DNS). Among these, LES offers high-fidelity predictions by resolving most of the energetic scales of turbulence, making it particularly suitable for accurate flow analysis. However, LES faces limitations in capturing near-wall phenomena, which require extremely fine meshes to resolve adequately.

Recent advances in deep learning, a branch of machine learning, have been applied to overcome some shortcomings of traditional LES models, especially in predicting subgrid-scale (SGS) terms. This thesis investigates the extension of existing multilayer perceptron (MLP) models from prior studies to 3D turbulent channel flow. The focus is on comparing network hyperparameters, feature sets, and training data to determine the most compact and efficient model capable of accurately reproducing SGS closure terms. Building on previous research at TU Delft on LES-VMM, this study emphasizes developing an effective network with a simple architecture that maintains high predictive accuracy while reducing computational complexity. ... Read more

Surfaces that induce slip at the wall have been shown to reduce skin friction drag in the turbulent regime. While superhydrophobic and liquid-infused surfaces are capable of drag reduction in turbulent hydrodynamic flows, neither would be feasible in aerodynamic applications due to the added roughness and the minuscule slip lengths created by the unfavourable viscosity ratio between the liquid and air. This thesis explores the novel idea of liquid-infused superhydrophobic surfaces (LISHySs). These surfaces use liquid droplets held in superhydrophobic features to induce slip at the air-liquid interface while reducing resistance to the motion of the liquid within the cavity.

Since no framework existed on the drag-reducing mechanism of such a surface, this was devised and the parameters of importance were identified. These were seen to be the area of the air-liquid interface, the exposed and submerged areas of the droplets in the cavities, and the degree of superhydrophobicity of the surface. Two types of LISHySs were consequently designed - one with spherical droplets and another with spanwise cylindrical droplets. These were produced by 3D printing surfaces with cavities and applying a superhydrophobic coating, into which water was infused.

Direct force measurements on the surface showed an increase in the drag coefficient (between 10% and 17%) with respect to a smooth reference surface. While PIV for the flow over this surface showed an increase in mean streamwise velocity and a decrease in Reynolds stress in the overlap layer, the high degree of reflection produced by the surface meant that meaningful near-wall data could not be obtained. Observations of the air-liquid interface demonstrated that the hypothesised steady rolling motion of liquid droplets occurred at low freestream velocities, but this was observed to become more chaotic at higher velocities. Pressure drag due to droplets acting as roughness elements, among other reasons, further contributed to the total drag produced.

While the results showed that the current implementation of LISHySs was unsuccessful, the lessons from this first attempt point to the directions in which future studies could be made. Cylindrical droplets oriented in the streamwise direction offer high potential due to the creation of streamwise slip. Advancements in the fields of material science and hydrophobicity would allow for surfaces with finer cavities that exhibit higher superhydrophobicity to be produced. Measures such as these would pave the way for the passive reduction of turbulent drag through the novel concept of LISHySs. ... Read more

As the space industry continues to grow rapidly, the development of reusable launch vehicles has become crucial in achieving cost-effective and sustainable access to space. Payload capacity optimisation has been at the forefront of this effort, leading to a renewed interest in hammerhead or bulbous payload fairing (PLF) configurations. These designs feature a larger diameter in the PLF than in the rest of the launch vehicle, enabling the same structure to be used for large payloads. However, the transonic regime poses unique challenges for these PLFs due to their susceptibility to flow separation and strong pressure fluctuations.

This master thesis investigates the influence of nose and boat-tail geometries on flow phenomena, particularly shock wave generation, around hammerhead configurations in transonic conditions. The research was conducted through three distinct experimental campaigns in the transonic wind tunnel TST-27 at the high-speed laboratory of TU Delft, employing schlieren, oil flow, and Particle Image Velocimetry (PIV) techniques.

The study revealed notable insights into the aerodynamics of the hammerhead PLF configurations. Boat-tails set at five and 15-degree angles were observed to broaden the range of shock wave oscillations, introducing an additional shock wave that could occasionally merge with existing ones. Moreover, the conic nose design induced higher shock wave oscillations, while the bi-conic nose introduced an extra shock wave compared to the conic and ogive noses. The study also found that altering the nose shape while keeping the boat-tail constant, or viceversa, resulted in similar effects on flow dynamics.

These findings underscore the critical role that nose and boat-tail geometries play in shaping the aerodynamic behaviour of hammerhead PLF configurations. The results have implications for such configurations’ design and stability considerations in transonic conditions. It was observed that raising the Mach number heightened shock wave oscillations and flow detachment. The angle of attack disrupted model symmetry, primarily impacting reattachment patterns. Furthermore, the conic nose exhibited greater unsteadiness due to oscillations in the shock waves compared to the other geometries. The study also provided detailed insights into shock wave spectral characteristics and identified potential influences of pressure wave oscillations on shock wave behaviour. The conclusions lead to recommend designs with bi-conic nose, and avoid boat-tail angles around 15 degrees.
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The need to reduce greenhouse gas emissions is emergent throughout every industrial sector worldwide. For aviation, this has opened a market gap for small scale electric aircraft configurations (e-V/STOL). These configurations are very diverse and they are classified based on how lift is obtained. However, their flight envelope consists of the same operational regimes which in turn, are very diverse compared with the ones of a traditional aircraft. This challenges the way wind tunnel testing was developed and urges the need for new testing techniques and wall correction methods. It is known for over a century that the flow inside a wind tunnel is not exactly the same as a free-air flow. This has to do with the existence of the wind tunnel test section boundaries which affect the flow over the model. Thus, the measured wind tunnel data deviate from the desired free-air performance measurements, except if the model is extremely small relative to the test section size. This is almost never satisfied due to physical scaling and structural considerations. Wall or boundary interference, is the quantification of those effects with the aim of correcting the wind tunnel performance measurements to their actual free-flight values. This is achieved by wall correction methods. These methods were developed in absence of high computing power and their purpose is to perform real-time adjustments to wind tunnel data. Consequently, they depend on simplified modeling assumptions, such as the linearized potential flow. As a result, these corrections might prove ineffective in generating reliable outcomes when dealing with non-linear boundary interference fields characterized by significant variations. For a given model, it is generally known that there is a limit at which wind tunnel test data are able to represent free-air conditions. This limit is called wind tunnel flow breakdown and it occurs when the model wake impinges on the boundaries of the wind tunnel test section, causing flow phenomena which would never occur in a free-flight situation. This phenomenon succeeds the non-linearities of boundary interference described above and is the worst case scenario for wind tunnel tests. The wall corrections completely fail to produce any valid outcome and the wind tunnel data are considered useless. This happens at low-speed/high-thrust conditions which are typical for a V/STOL aircraft during transition from hover to forward flight. Thus, powered wind tunnel testing at such conditions poses challenges in terms of data representation to an interference-free, free-air situation. Such conditions may also be encountered when the aircraft is operating in-ground vicinity and thus, the interference field for those cases is of interest. In the latter case, it is expected that the magnitude of the corrections and the wind tunnel flow breakdown limits would be significantly lower. This Thesis aims to investigate wall interference effects between the rotor and the wind tunnel test section boundaries at low-speed/high-thrust conditions, establish the flow breakdown limits for a given propeller and test section and to gain insight into the effect that the bottom wall of the test section exhibits to the boundary interference field. This is realized by conducting wind tunnel tests on two different fixed-pitch propellers with varying incidence angles in both a closed-wall and a 3/4 open-jet test section of the NLR Aeroacoustic Wind Tunnel. Static wall pressure measurements are acquired and the flow-field in the vicinity of the propeller is quantified using large-scale, Tomographic Particle Tracking Velocimetry (PTV) with the use of helium filled soap bubbles (HFSB) as tracers. Finally, the interference-free performance data for these propellers are obtained from the second test campaign performed in the industrial Low-Speed-Wind Tunnel (LST) of the DNW. The global flow topology of the wall-bounded tests in the AWT-closed test section is presented by means of wall pressure measurements and PTV. It was found that a distinct rise of the pressure coefficient leading to a maximum peak, corresponds to wake impingement. This firstly occurs on the advancing side of the rotor due to the larger vortex which is deflected further downward than the one on retreating side. A flow breakdown criterion which makes use of the static pressure distribution on the bottom wall of the test section is introduced. It aligns well with existing benchmarks available in literature for closedwall test sections. This criterion can be applied to any test section consisting of a solid lower wall, as long as the distribution of static pressure is monitored. On the other hand, the comparison between the LST and AWT-closed test section data do not show any signs of flow breakdown. This could prompt additional questions regarding the necessity of an even larger test section for obtaining the interference free propeller performance data. Various wall correction methods are utilized in order to evaluate their applicability for powered wind tunnel testing and to assess the impact of wall interference. It is shown that the developed Two-VariableMethod (TVM) does not provide satisfying results with respect to lift interference, when compared with the Heyson method. Nevertheless, blockage predictions follow the expected trends even for lifting cases, at least for the higher advance ratios where no flow impingement is present. For a simple non-lifting case, this method is deemed more reliable since its underlying assumptions are satisfied to a greater extend and its blockage predictions are also in good agreement with the Glauert’s method. For the purely lifting angle of attack, the comparison between the two test campaigns is not satisfactory both for in-ground and out-of-ground effect. This discrepancy arises from the fact that the predictions of all wall correction methods applied do not abide by the trend that is discernible based on the deviations of the bounded (AWT) to the unbounded (LST) data. That is mainly attributed to the load measurement system and to the possibly stalling conditions at the higher wind tunnel speeds. For the 0◦ incidence angle, there was a very good agreement between the AWT-closed test section and the LST data for both flight conditions (OGE, IGE). This was validated by both applied correction procedures (Glauert’s method for off-center propellers & Two-Variable-Method). Even in that case, the AWT-3/4 open-jet results were not in good agreement implying effects which may not only be attributed to boundary interference. ... Read more

Surface pressure measurements are a crucial part of aerodynamic/hydrodynamic analysis and are essential when studying complex flow phenomena. Although pressure-sensitive paint (PSP) has become established as a non-intrusive field measurement technique for high-speed applications in air, a comparable method does not yet exist for usage underwater.... ... Read more

Unmanned aerial vehicles have proliferated in the last few decades, with applications that include military, commercial and recreational. Their size and typical flight velocities are characterized by moderate Reynolds numbers O(10^4-10^5). For such conditions, boundary layers can remain laminar and therefore are highly susceptible to separation and the generation of laminar separation bubbles (LSBs), when compared to more conventional aircraft. Extensive research has been done to study the influence of Reynolds number, angle of attack, sweep angle or freestream turbulence level on the nature of LSBs. However, the study of LSBs subject to unsteadiness is rather limited. This is especially relevant for small aircraft that typically fly in gusty environments such as cities. The problem is aggravated by the recent shift toward composite manufacturing, which allows more efficient high-aspect-ratio configurations, but that deform considerably more when subjected to unsteady loads.
The LSB that forms on the suction side of a modified NACA 64_3-618 airfoil at a chord-based Reynolds number of Re = 200k is studied in a series of wind tunnel experiments conducted at The University of Arizona. Three different flow measurement techniques are considered in the experiments to identify the bubble: surface pressure measurements, Particle Image Velocimetry and Infrared Thermography. The capabilities of the three techniques are first explored in a static characterization of the LSB over a range of angles of attack. For the conditions tested, excellent agreement between the techniques is obtained, showing an upstream shift of the bubble with increasing incidence. For the study of static LSBs, the infrared approach is superior, given its higher spatial resolution and experimental simplicity.
The complexity is then increased to study the influence of aerodynamic unsteadiness on the bubble. For this purpose, two different types of structural motion are imposed on the wind tunnel model. A first experiment considers a pitching-type motion, with reduced frequencies up to k = 0.25. While surface pressure measurements and PIV are not heavily affected by the change in experimental conditions, the infrared approach becomes limited by the thermal response of the surface. To overcome this limitation, an extension of the recently proposed Differential Infrared Thermography (DIT) method is considered. Even so, the unsteady behaviour of the bubble can be only partially detected with this method. All-three techniques considered indicate a hysteresis in bubble location between the pitch up and pitch down parts of the motion, caused by the effect of the aerodynamic unsteadiness on the adverse pressure gradient.
The second type of structural motion studied consists of a sinusoidal plunge, with an amplitude of h = 6% of the airfoil chord and a reduced frequency of k = 0.67. The surface pressure measurements and PIV still capture a hysteresis in bubble location along the cycle, expressed in terms of the effective angle of attack induced by the plunging motion. However, due to the increased frequency of the motion, the thermal response of the surface reduces and the infrared approach fails to detect the unsteady bubble.  ... Read more

The technical advancements achieved during the last decade have permitted the implementation of low-fidelity numerical methods to analyze simplified aeroelastic problems. However, the available computational power is still incapable of dealing with more realistic events involving complex structural dynamics and flow regimes. For this reason, experimental wind tunnel campaigns are still required to characterize relevant fluid-structure interaction events and to assist in the development of more accurate computational models. Nevertheless, conventional measurement systems present several constraints that limit the complete description of the phenomena being analyzed. This thesis is part of the research effort to define alternatives capable of providing full-field and simultaneous information of both flow and structural quantities in a non-intrusive way.

A wind tunnel experimental campaign has been developed in the Open Jet Facility to evaluate the ability of Lagrangian Particle Tracking (LPT) to characterize aerodynamic loads acting on large-scale flexible bodies. Three different load reconstruction approaches have been considered for this purpose: integral momentum conservation equation in its classical formulation, integral momentum conservation equation with an alternative formulation based on vorticity, and Kutta-Joukowski theorem. The accuracy and precision of each of the forenamed procedures have been assessed in both steady and unsteady inflow conditions taking force balance measurements as a reference.

The results of this thesis have proven the capabilities of LPT to determine the aerodynamic loads in aeroelastic wind tunnel investigations with acceptable accuracy and precision. Finally, it has also been demonstrated that the range of information provided by these techniques is superior to that achievable with conventional measurement devices, which improves the overall characterization of the fluid-structure interaction phenomena. ... Read more

Supersonic flows cause thin panels to flutter, which is characterized by high-amplitude self-sustained oscillations, increasing the risk of fatigue failure. Flutter is known to be exacerbated when a shock wave impinges on the panel, creating a shock wave/boundary-layer interaction (SWBLI) which promotes separation of flow and leads to increased aerodynamic and thermal loading on the panel. This novel fluid-structure interaction (FSI), known as shock-induced panel flutter, poses a risk to the structural integrity of components used in high-speed aerial vehicles, such as rocket nozzles and supersonic engine inlets.

Due to the complex coupled interaction between flow and structural dynamics in flutter, experiments provide a better option to investigate the underlying physical mechanisms, rather than computational studies, which usually have to compromise between computational cost and resultant accuracy. The ST-15 supersonic wind tunnel facility at TU Delft is employed to conduct experimental measurements of shock-induced panel flutter, using high-speed Schlieren imaging to capture flow structures and stereographic Digital Image Correlation (DIC) to record panel displacements in separate campaigns. Tests are done at Mach 2 with fully-clamped thin panels, and the flutter response is excited using different shock strengths and impingement locations.

In both campaigns, accelerometers are used to measure spurious vibrations around the wind tunnel test section. This helps reveal the existence of external vibrations inherent to the facility, which are also found to drive the frequency of the panel flutter: at 760-770 Hz without an impinging shock, and 605-640 Hz with an impinging shock. The latter frequency is detected in both - shock motion and panel displacements – which establishes coupling between flow and structure despite measurements being non-simultaneous. Changing the shock impingement location does not have a significant effect on the degree of flow separation caused over the thin fluttering panel, which is always higher than the separation on a rigid plate at the same shock strength, thus proving that fluttering panels are not viable means of shock-induced separation control. A stronger impinging shock produces increased shock-induced flow separation but results in less energetic panel flutter, which is attributed to the higher post-shock pressure rise suppressing the panel. Flutter is found to be most energetic, and consequently, the panel is most susceptible to fatigue failure, when the shock impinges at 60% of the panel length. ... Read more