Surface imperfections on laminar aircraft surfaces, such as panel joints and discontinuities, can strongly modify boundary-layer stability. Sharp-edged rectangular bumps (SERBs) combine two widely studied surface features, the Forward-Facing Step (FFS) and a Backward-Facing Step (BFS), and provide a useful geometry for examining how step-induced mean-flow distortion and separation affect the development of Tollmien-Schlichting (TS) waves. However, previous studies on similar bump shapes have mostly considered a single incoming TS-wave amplitude, leaving the dependence of the SERB-induced boundary-layer response on incoming TS-wave amplitude insufficiently resolved.

To investigate how the boundary-layer response to a SERB depends on the amplitude of an incoming forced TS wave, this work combined numerical stability analysis with experiments performed on an unswept flat plate in the low-turbulence environment of the anechoic wind tunnel (A-tunnel) at TU Delft. Dielectric-Barrier Discharge (DBD) plasma actuators were used to force single-frequency TS-wave disturbances while allowing for easy amplitude variation. Surface microphones and Hot-Wire Anemometry (HWA) measurement systems were used to collect data to analyse the harmonic content of the disturbance, its wall-normal structure at the forcing frequency, and its downstream growth, comparing across the clean flat plate with and without the SERB installed over a range of forcing amplitudes. Independent monitoring of forcing consistency was shown to be important in long-duration plasma-forced experiments, where maintaining a consistent forcing amplitude is required. Simultaneous microphone measurements were necessary to distinguish genuine flow-induced amplification from transient variations in actuator output.

The results show that the influence of the SERB is strongly region-dependent: the forced TS-wave behaves like the clean case upstream of the bump and undergoes local amplification in the regions immediately upstream of the FFS and in the region just downstream of the FFS lip, a reduction in growth rate and stabilisation over the bump, and greater amplification downstream of the BFS. Greater forcing amplitudes are associated with a broader redistribution of spectral energy in the wall-normal direction and an earlier loss of a distinct shear-layer instability signature introduced in the recirculation region behind the BFS. The SERB, therefore, does not behave as a simple amplifier of TS waves, but instead produces a region-dependent response, with local amplification ahead of the FFS, reduced growth over the bump, and stronger re-amplification downstream of the BFS, where the response becomes increasingly sensitive to the amplitude of the incoming disturbance.

This thesis investigates how surface deformations can delay the transition from laminar to turbulent flow in aircraft boundary layers, thereby reducing drag and improving fuel efficiency. This is important because aviation contributes significantly to greenhouse gas emissions, and improving the lift-to-drag ratio is one of the most effective ways to make aircraft more sustainable.

The focus is on the early stages of boundary layer transition, particularly the growth of stationary crossflow instabilities in swept-wing boundary layers. These instabilities play a major role in triggering turbulence on swept wings. While experiments and Direct Numerical Simulations (DNS) can study this process accurately, they are computationally expensive. Therefore, this thesis uses flow stability analysis methods, which are much faster.

Three stability-analysis approaches are discussed. The classical Orr–Sommerfeld method is computationally efficient but assumes a locally parallel flow and only works for small perturbations. The Parabolized Stability Equations (PSE) improve on this by including streamwise development and nonlinear interactions, but they are only valid for slowly varying flows. The Harmonic Navier–Stokes (HNS) equations retain all streamwise derivatives and can therefore handle strongly non-parallel flows, although at a higher computational cost than PSE.

To exploit these advantages, a new computational framework, the Delft Harmonic Navier–Stokes Solver (DeHNSSo), is developed. DeHNSSo can analyse the effect of both smooth and sharp surface deformations, such as humps and steps, on boundary layer instabilities. The solver uses a Fourier-based representation of perturbations, spectral discretisation in the wall-normal direction, and finite differences in the streamwise direction. Nonlinear interactions between perturbation modes are included iteratively.

The framework is validated using several standard instability cases, including Tollmien–Schlichting waves in a Blasius boundary layer and stationary crossflow instabilities in a swept-wing boundary layer. In all cases, DeHNSSo agrees closely with DNS and with other stability methods such as PSE and Adaptive Harmonic Linearised Navier–Stokes.

The main application of the solver is to investigate the effect of a shallow smooth surface hump on crossflow instabilities. The hump creates a local region of reversed crossflow without causing flow separation. Away from the hump, the boundary layer quickly returns to its original state.

For small perturbation amplitudes, the hump reduces the growth of crossflow instabilities over a large downstream region. Although there is some local destabilisation near the hump, the overall effect is stabilising because the perturbation shape is altered in a way that weakens the lift-up mechanism responsible for instability growth.

At larger perturbation amplitudes, however, the hump becomes less effective. A second unstable mode appears near the wall and transfers energy to the main instability. This can lead to locally increased disturbance amplitudes and earlier quasi-saturation, potentially accelerating transition.

The results suggest that smooth surface humps are most effective when the incoming disturbances are still small and approximately linear. Under those conditions, they can delay transition and reduce drag. The study therefore provides a promising basis for using optimised surface humps on aircraft wings. Future work could further optimise hump shape and arrangement and improve computational efficiency by coupling the HNS approach with faster methods such as nonlinear PSE.

Swept wings of commercial passenger jets experience skin friction drag from the turbulent boundary layers over them. Unique to such wings is the turbulent transition induced by a crossflow instability, where the laminar boundary layer diverges from the inviscid airflow due to competing forces. Dielectric Barrier Discharge (DBD) actuators have minimised this divergence by applying a uniformspanwise plasma body force against the boundary layer, and delayed transition. Yet, the unsteadiness of the AC plasma discharge can induce early transition, and weak body forces limit their control effectiveness at high Reynolds numbers. Therefore, thisMasters thesis characterised the effects of Extended DBD actuators, where the presence of a positive DC electrode to maximise body forces and minimise unsteady disturbances in the plasmawas studied. Promising results were seen in the form of increased body forces and decreased Turbulent Kinetic Energy addition to the flow, compared to a traditional DBD actuator.

This thesis presents a comprehensive numerical investigation of wall interference effects in transonic wind tunnel testing of swept-wing configurations, by the example of the Cryogenic Ludwieg Tube Göttingen (KRG) at the German Aerospace Center (DLR). By employing Reynolds-Averaged Navier–Stokes (RANS) simulations using the DLR TAU code and the Spalart–Allmaras turbulence model, the study quantifies how test section walls affect the spanwise flow homogeneity around a swept wing at high Reynolds numbers. A detailed analysis was conducted to isolate the influence of key flow parameters, like angle of attack, sweep
angle, chord length, and Mach number, on wall-induced distortions.

The simulations reveal that vertical tunnel walls introduce significant spanwise pressure gradients, especially at higher sweep angles and larger chord lengths, i.e., aspect ratios. These effects are mitigated through a mesh deformation approach based on Radial Basis Functions, which enables the adaptive reshaping of tunnel walls to align with streamlines from an idealized, unbounded flowfield. The implementation of these deformed geometries resulted in substantial improvements in spanwise flow uniformity, reducing the root mean square errors in the pressure coefficient distribution by nearly 50% in critical test section regions. Not only is it possible to reduce spanwise gradients in the flowfield, but also to reduce the difference to infinitely swept conditions almost over the entire width of the test section.

Beyond identifying optimal strategies to reduce the influence of the walls on the flowfield, the thesis validates its numerical findings against existing experimental and numerical data. The proposed methodology offers a valuable toolset for pre-test planning and supports the development of more representative aerodynamic experiments, especially in the context of laminar-to-turbulent transition studies. These contributions provide a scientifically robust foundation for minimizing wall interference in transonic testing and support the advancement of research on laminar-to-turbulent transition.

Hot-wire anemometry (HWA) remains an essential diagnostic tool for high–frequency velocity measurements in aerodynamic research, yet its practical accuracy strongly depends on robust temperature correction and reliable multi–component probe calibration. Despite the maturity of the technique, the literature provides no consensus on how best to correct for fluid–temperature variations or how to calibrate X-wire probes with consistent accuracy across wide velocity and angular ranges. The situation is further complicated at the TU Delft Low-Speed Wind Tunnel Laboratory (LSL), where the absence of an in-situ multi-wire calibration system has made efficient and repeatable probe preparation difficult. This thesis addresses these issues by developing a new automated calibration framework for multi-wire CTA probes and performing a comprehensive evaluation of existing temperature-correction and X-wire calibration schemes.
The first part of the work concerns the separation of velocity and temperature effects in CTA measurements. Six temperature-correction models—ranging from simple Bearman [1]-type scaling ignoring the change in fluid properties with temperature, to more complex formulations incorporating fluid-property variations and the Collis & Williams factor [2] are applied to four independent datasets from literature, covering a wide range of velocities and temperature conditions. Each model is assessed by collapsing data acquired at varying temperatures onto a single fourth-degree polynomial calibration curve to quantify the residual error via normalized root mean square percentage error. Across all datasets, simple temperature corrections perform surprisingly well. Relying on resistance-based wire-temperature estimation suffers from uncertainty, with the best success when calculating the fluid properties at the film temperature and applying the Collis & Williams correction. Allowing the wire temperature to act as an optimization parameter can improve the fit of the curves, confirming earlier findings that resistance-derived temperatures can introduce systematic error.
The second part of the thesis focuses on the calibration of X-wires. A new miniature, high-precision, in-situ yaw calibrator is designed and implemented. The device integrates a compact high-torque servo motor and a magnetic rotary encoder with high accuracy, assembled using high-precision alignment procedures. With the calibration and controlled through a LabView interface, and the position feedback and servo control integrated via an Arduino implementation, the system enables automated yaw sweeps. Its compact form factor allows installation directly inside restricted windtunnel test sections.
Using this system, a dense X-wire calibration dataset was collected, spanning velocities from 5 to 30ms−1 in 1ms−1 increments and yaw angles from −40◦ to 40◦ in 1◦ steps. The dataset enabled comparison of calibration schemes, including interpolation-based indirect methods (Lueptow [3] and Tropea [4] variants), and direct polynomial surface

This thesis presents a theoretical and numerical investigation of how a surface forward-facing step alters the stability and transition mechanisms of a laminar incompressible swept-wing boundary layer dominated by stationary crossflow instabilities (CFI). The results elucidate the processes that cause significant transition advancement due to the step and, in turn, challenge the classic paradigm in fluid mechanics that rapid, localised surface-geometry variations are universally detrimental to laminar flow. Through Direct Numerical Simulations (DNS) and new theoretical modelling, this thesis shows that, under specific conditions, a forwardfacing step stabilises pre-existing stationary crossflow vortices and thereby delays laminar-turbulent transition...

The work presented in this dissertation focuses on the developed and implementation of real-time control techniques for turbulent wall-bounded flows, with the aim of achieving skin-friction drag reduction. After an initial control system was developed that utilized instantaneous wall-shear stress fluctuations, wall-pressure fluctuations were subsequently used as the input quantity to the real-time flow control systems considered in this dissertation. Furthermore, the control algorithm complexity was escalated from a relatively simple feedforward opposition control logic to an adaptive control strategy. The research presented in this dissertation is fully experimental in nature. Experimental activities were conducted in two main facilities. The bulk of testing was performed in the W-Tunnel at Delft University of Technology: an open-return wind tunnel, where a modular test section was integrated to perform experiments on zero-pressure–gradient turbulent boundary layer flows. A subset of measurements were conducted at the Center for International Cooperation in Long-Pipe Experiments (CICLoPE) at Bologna University, in Italy. Here, simultaneous measurements were conducted of velocity fluctuations in the logarithmic region and wall-pressure.

A stochastic spectral correlation analysis between wall-pressure fluctuations and velocity fluctuations in the logarithmic of a turbulent pipe flow reveal Reynolds-number– independence of the wall-pressure linear coherence spectrum. This is a first-of-its-kind result, hinting at the feasibility of scaling an input sensing strategy based on wall-pressure fluctuations from a low-Reynolds-number environment to operational engineering conditions.

An initial controller based on wall-shear stress fluctuations was developed to target drag-producing large-scale structures in the logarithmic region. The flow response was measured in terms of both the statistical (and spectral) response of the TBL flow to realtime control and in terms of the effect the control has not only on the friction coefficient, but also on the integral measures. This analysis revealed three main findings: (1) an attenuation of energy at streamwise wavelengths characteristic of large-scale motions, (2) a decrease in skin-friction and (3) an attenuation of the statistical integral measures of skin-friction (i.e. bulk production and FIK terms).

A similar control architecture was also implemented that employed wall-pressure (and wall-pressure–squared) as the input quantity. It was found that the wall-pressure–squared term improves the accuracy of an estimator enabling the prediction of off-the-wall velocity fluctuations from a wall-based position. Furthermore, its inclusion is essential, given that the linear term does not retain sufficient coherence over the relatively large streamwise extent separating input and actuation locations.

The final controller that was developed in the context of this dissertation is an adaptive one, relying on the Filtered-X Least Means Squares (Fx-LMS) algorithm. This strategy does not rely on a-priori system identification, as was the case for the previous two control strategies mentioned above. Instead, it automatically identifies the coefficients of the transfer functions relating input to output in the controller. For this study, this algorithm was deployed both to a flow case that was strongly modulated by cylinder vortex shedding and to a fully broadband turbulent boundary layer flow. In the former case, the controller readily identified the shedding frequency as the control target. For the latter, the controller converges to a situation where the large-scales were targeted and their intensity successfully attenuated.

This thesis project undertakes a comprehensive numerical analysis of aerodynamic cooling ducts in the context of fuel-cell powered aircraft, aiming to enhance the efficiency and performance of these innovative propulsion systems. The use of fuel cells in aviation presents a paradigm shift towards sustainable and environmentally friendly air travel. However, the integration of fuel cells introduces new challenges, particularly in managing the excess heat generated during operation. Aerodynamic cooling ducts play a crucial role in dissipating this heat while minimising aerodynamic drag.

The numerical analysis involves the application of aerodynamics, thermal management and performance techniques to model and simulate the complex airflow within the cooling ducts. Parameters such as duct geometry, airflow properties, and heat transfer rates are systematically investigated to optimise the cooling process. The study also explores the interaction between the cooling ducts and the overall performance of the aircraft, considering the impact on drag and fuel efficiency.

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%.

Maintaining laminar flow on large swept surfaces of subsonic transport aircraft, i.e. the wings and the stabilisers, is currently posing a considerable challenge for aerodynamic design. Improving the efficiency of aircraft by delaying or removing the laminar-to-turbulent transition process over the wing and tail parts can substantially reduce contaminant emissions. The dominant flow instability causing laminar-turbulent transition of swept-wing flow is the so-called crossflow instability (CFI). Ongoing research at TU Delft has shown potential to delay transition by use of passive mechanisms. As such, a framework has been designed to numerically compute crossflow development and transition to turbulence on swept wings.

Through the use of experimental data acquired in wind-tunnel measurements at TU Delft, the CFI development and transition process on swept wings has been modelled numerically by means of Direct Numerical Simulation (DNS). Based on a DNS laminar flow field generated from the pressure distribution along the model surface, a numerical primary CFI mode in good agreement with the experiment was obtained through Non-linear Parabolized Stability Equations (NPSE). Following this steady flow field analysis, the simulation was made unsteady by the implementation of numerical free-stream turbulence. This novel method resulted in unprecedented modelling of the receptivity mechanisms of transition in three-dimensional crossflow cases, overcoming ad-hoc treatments.

Both experimental and numerical flow fields indicated a Type-I dominant secondary CFI (i.e. KH-type response in the laterally inclined shear layer of the stationary crossflow vortex), which consequently carries the formation of near-wall hairpins and ultimately turbulence. Crossflow vortex frequency content also agrees well in the low-frequency band (450 Hz ≤ f ≤ 3000 Hz), whilst the numerical high-frequency content (3500 Hz ≤ f ≤ 9000 Hz) does show a distinct delay in amplitude growth throughout the majority of the transition region.

Contradicting the promising qualitative analysis of the free-stream turbulence methodology, this discrepancy in the frequency spectrum indicates a major shortcoming in the numerical setup, which was shown to be biased towards introducing more low-frequency disturbances at the inflow boundary.

An experimental investigation is presented in which an array of pulsed jet actuators is used to control a turbulent separation bubble formed on a curved backward facing ramp. The array is positioned upstream of detachment and consists of wall-normal high aspect ratio skewed rectangular jets which generate streamwise vortices in the boundary layer increasing momentum transfer and delaying separation. While similar systems have shown promise in previous research, this work considers a pressure-induced separation of a relatively high Reynolds number (Re_τ=4600) turbulent boundary layer (TBL), where the large turbulent structures of the separating BL are of similar scale and magnitude as those generated by actuation and significantly affect the dynamics of detachment.

Both steady blowing and periodic pulsing actuation strategies are tested and compared. Preliminary jet velocity and pulsing frequency sweeps are carried out to identify optimal actuator operating parameters, relying on wall static pressure measurements to evaluate control effectiveness. Select cases of interest are then investigated using two-dimensional two-component particle image velocimetry and compared against the uncontrolled baseline which is characterized using PIV and hot wire anemometry. Additional PIV-derived metrics are utilized to assess system performance.

For steady blowing, a jet-to-crossflow velocity ratio VR>1 was required to produce a separation delay, while diminishing improvements in control effect with increasing jet velocity started at VR=1.6 (actuation momentum ratio of C_μ=2.3%). This nominal velocity ratio was adopted for all further investigation. The actuator was found to produce alternating strong and weak downwash regions in the TBL resulting in an artificial sweep/ejection pattern at detachment. Periodic forcing with the same nominal velocity ratio was able to achieve better or comparable results to steady actuation, while requiring less input momentum (C_μ=1.2-1.8%). The optimum actuation frequency was determined to be the natural frequency of the uncontrolled bubble, with the performance of higher frequency actuation tending towards steady blowing levels. As shown by an analysis of flow dynamics based on phase-averaged PIV velocity fields, actuation at the bubble time scales produces significant flow oscillation in phase with actuation. This resonant behaviour results in transient high momentum sweeps between actuation pulses that boost actuator performance, achieving double the performance benefit afforded by steady actuation according to multiple metrics. In comparison, actuation at time scales multiple times shorter than that of the bubble produces a quasi-steady flow and performance comparable to that of steady actuation.

Additionally, a novel alternating actuation strategy is tested, in which the period of active blowing is composed of high frequency alternation between two inverted actuator rows. This aimed to produce a quasi-2D periodic control effect using 3D actuators, which Squire's theorem suggests could excite the separated shear layer instability more than conventional 3D perturbation. While high frequency alternation did achieve a quasi-2D effect, it also prevented the sweep/ejection pattern characteristic of 3D perturbation from forming, thus significantly limiting the actuator performance.

The transition behaviour of a boundary layer with zero pressure gradient in a low-subsonic freestream is examined for a set of wall temperatures below the freestream temperature. Solutions of the boundary layer provided by both an incompressible and a compressible numerical flow solver are compared to wind tunnel measurements on a uniformly cooled section of a flat plate. The compressible solver features a temperature dependence of the fluid properties, which proves to be crucial for modeling the stability of a thermal boundary layer. Thermal images of the surface show that the spanwise-averaged transition front moves downstream with decreasing wall-to-freestream temperature ratio. At the streamwise station where no frictional heating due to turbulent structures is observed, T-S waves are detected in the measured velocity fluctuation profiles. The experimental data show good correspondence with the numerical linear stability predictions here. Further downstream, where streaks of transition occur, T-S waves can no longer be distinguished, and it is likely that non-linear effects have overtaken the boundary layer flow here. The absolute amplitude and the amplitude growth rate of the velocity perturbations both decrease with the wall temperature in the region where T-S waves are seen, and even more in the non-linear region. It is concluded that uniform surface-cooling stabilizes the boundary layer on a flat plate in an incompressible freestream, but a variable-fluid property solver is required to model the stability characteristics.

The research presented in this thesis focuses on the receptivity to surface roughness of swept wing boundary layers dominated by crossflow instabilities (CFI), providing insights into how surface roughness can be used to passively control the developing instabilities. Discrete roughness elements (DRE) arrays and distributed randomized roughness patches (DRP) are employed to investigate the physical phenomena governing receptivity and their impact on CFI onset. The supporting data combine numerical solutions of linear and non-linear stability theory with advanced experimental flow diagnostics.

This booklet is divided into three main parts. The first part investigates the flow mechanisms dominating the receptivity of stationary CFI to the amplitude and location of DRE arrays. The relation between the external forcing configuration and the initial instability amplitude is investigated, along with scaling principles allowing for the up-scaled reproduction of the swept wing leading-edge configurations, which provide experimentally observable configuration.

The second part of this research explores the stationary CFI receptivity to specific up-scaled roughness configurations, including both isolated discrete roughness elements and DRE arrays. These roughness elements are applied at relatively downstream chord locations to enhance the experimental resolution of the near-roughness flow field.

The isolated discrete roughness elements ensure strong boundary layer forcing, which helps to outline the relation between the near-element instability onset and the rapid transitional process. In contrast, the applied DRE arrays configurations provide boundary layers dominated by the development of CFI. In such scenarios, high-magnification tomographic particle tracking velocimetry identifies the dominant near-element stationary instabilities precursor to CFI. Specifically, the presence of transient growth and decay mechanisms in the near-roughness flow region is outlined, exploring their role in the receptivity process and in the CFI onset. This investigation results in the first conceptual map describing the receptivity of swept-wing boundary layers to a wide range of DRE array amplitudes.

Lastly, the acquired knowledge of the near-element flow topology is employed in the final part of this work to develop a passive laminar flow control technique for stationary CFI cancellation. This technique is based on the destructive interference of the velocity disturbances introduced by a streamwise series of optimally arranged DRE arrays. The performed measurements confirm a reduction in the developing CFI amplitude accompanied by a delay of the boundary layer transition. The compatibility of the proposed technique with the control of CFI developing in a realistic free-flight scenario is as well investigated.

When designing an airborne wind energy system, it is necessary to be able to estimate the traction force that the kite produces as a function of its flight trajectory. Being a flexible structure, the geometry of a soft kite depends on its aerodynamic loading and vice versa, which forms a complex fluid-structure interaction (FSI) problem.
Currently, kite design is usually done on an experimental basis since no model meets the requirements of being both accurate and fast.

In this project, an FSI methodology is developed to study the steady-state aerodynamic performance of leading-edge inflatable (LEI) kites by coupling two fast and simple models.

On the structural part, the deformations are calculated with a particle system model, based on the assumption that the shape of the kite can be modelled using a wireframe wing model represented by the bridle line attachment points, whose coordinate changes are modelled using a bridle
line system model and canopy billowing relations.

On the aerodynamic side, the load distribution is calculated with a 3D nonlinear vortex step method, coupled with 2D polars obtained with a correlation model derived from Reynolds averaged Navier-Stokes (RANS) analysis, to account for viscous effects and flow separation. Furthermore, with the 2D correlation model it is possible to consider changes in the thickness and the camber of each section. Based on 2D thin airfoil theory, the three-quarter chord point is used to determine the magnitude of the forces, and the one-quarter chord point is used to determine the direction of these forces.
Moreover, the model developed for LEI kites can consider canopy billowing and variations in kite and airfoil geometry while proving robust and inexpensive.
This model has been validated with several geometries and a RANS analysis of the LEI kite, showing great accuracy for pre-stall angles of attack.

The coupling of these two models results in a fast aeroelastic model of LEI kites capable of predicting the steady-state deformations and aerodynamic forces on the kite for the range of actuation settings and inflow conditions expected during a normal pumping cycle. Furthermore, the results show that the deformations follow the same trends as the results from the photogrammetry analysis and that, by taking into account the deformations that the kite undergoes, the aerodynamic forces more closely resemble experimental data.

Current manufacturing techniques in the aviation industry result in a number of two-dimensional surface irregularities (e.g. panel joints, seals, seams) which can lead to an increase in skin friction due to premature boundary layer transition. Panel joints are commonly modeled in the form of two-dimensional steps. Prior studies have shown that Backward-Facing Steps (BFS) promote transition earlier than Forward-Facing Steps (FFS). Therefore, in the design of laminar flow components, FFS are preferred over BFS. However, the understanding of which mechanisms are responsible for the larger growth experienced by Tollmien-Schlichting (TS) waves in the presence of FFS remains unknown. Prior experimental works focused on parametric studies on transition location with limited measurements close to the step. In addition, studies using Direct Numerical Simulations (DNS) assume two-dimensional flow and consequently do not capture transition location. All of this makes comparison between existing experimental and numerical data rather cumbersome, hindering the problem understanding.

In light of this, the present study aims to close-examine the TS waves at the step to identify which are the relevant mechanisms that modify their growth and move transition upstream. To do so, this work presents an experimental and numerical investigation jointly conducted by TU Delft and the German Aerospace Center (DLR) on Tollmien-Schlichting (TS) waves interaction with a Forward-Facing Step (FFS). Experiments are conducted at the TU Delft low-turbulence anechoic wind tunnel (A-tunnel) on an unswept flat plate model. Single-frequency disturbances are introduced using controlled acoustic excitation. The temporal response of the flow in the vicinity of the step is measured using Hot-Wire Anemometry (HWA). In addition, the global effect of the step on laminar-turbulent transition is captured using Infrared Thermography (IR). Two-dimensional (2D) Direct Numerical Simulations (DNS) performed at DLR provide detailed information at the step. Experimental and numerical comparison is performed in subcritical step conditions. At larger step heights only experimental data is provided.

Experimental and DNS results in clean and subcritical step conditions present very good agreement. Both methods predict large distortion of the TS wave downstream of the step, where DNS results present different growth trends between streamwise and wall-normal components of the fundamental mode. Furthermore, while upstream of the step the TS waves exhibit exponential growth, downstream of it they present a complex growth behavior followed by regions where the perturbation energy production term changes sign along the streamwise direction. Finally, regions of highly negative and positive production seem to correlate with the tilting of TS waves in and against the mean shear direction, respectively. These findings point towards the presence of different growth mechanisms triggered by the step which could modify the level of amplification of disturbances far downstream.

Laminar-to-turbulent boundary layer transition on the wings and stabilizers of aircraft leads to a large increase in the skin-friction drag they experience during flight. This has motivated prolific research into understanding and delaying the transition process, in an effort to improve the economic and ecological impact of air travel. The swept wing poses a particular challenge to this, as the transition scenario in flight is typically governed by the growth and breakdown of stationary crossflow instabilities (CFI). These are destabilized by flow conditions that would suppress instabilities on straight wings, such as Tollmien–Schlichting waves, complicating the use of laminar flow control techniques. Previous research has shown stationary CFI to be highly receptive to surface roughness, with the use of discrete roughness elements (DRE) having emerged as a forcing/control strategy. However, the mechanisms behind which DRE condition the onset of CFI in a swept wing boundary layer are still a topic of ongoing investigations.

The research presented in this Master thesis aims to further characterize the relation between the DRE forcing configuration and the onset of swept wing boundary layer instabilities, through an investigation of its wake flowfield. The DRE wake has previously been observed to be highly non-modal in nature, limiting the use of modal linear stability theory in predicting stationary disturbance evolution in the element vicinity. To address this, a linear, non-modal, parabolized stability framework is derived, capable of simulating transient growth within the DRE wake region. The numerical framework is initialized with experimental DRE wake data from a previous study, for a DRE array of critical forcing height. The predicted stability characteristics qualitatively agree with experiment, with the numerical solver successfully evolving the stationary DRE wake structure into stationary CFI downstream of the array. However, a quantitative match with experiment could not be attained. The results were also noted to be sensitive to the choice of initial conditions, as well as to the numerical discretization.

In addition to the stationary numerical simulations, hot-wire anemometry measurements are conducted at the TU Delft low turbulence wind tunnel, providing a first-time experimental characterization of the evolution of unsteady disturbances in the DRE wake for a swept wing boundary layer. Regions of high fluctuation intensity are observed to be localized around the top and sides of each element in the DRE array, with their spectral characteristics associated to circular cylinder shedding. For a critical DRE forcing, the evolution of the wake flow is largely governed by the stationary structures, although unsteady disturbance energy is observed to undergo a brief period of transient energy amplification. A super-critical DRE forcing introduces strongly tonal fluctuations into the DRE wake, whose amplification is enhanced by interaction between the stationary structures. The high fluctuation levels are sustained and grow to non-linear amplitudes, inhibiting the relaxation of steady disturbance energy. This leads to rapid, local turbulent breakdown in the vicinity of the elements.

The flow of air over a swept wing initially starts in a smooth laminar state (referred to as base flow) and entrains disturbances which subsequently grow and transition the flow to a chaotic, turbulent state. Active efforts have been made to study and control these disturbances, which manifest as stationary crossflow modes. Excrescences in the form of a forward-facing step (FFS) impose a base flow deformation in the form of a rapid near-wall pressure change, flow separation, and strong upwash. This modifies the behaviour of stationary crossflow instability and subsequently leads to an upstream or downstream shift of transition location depending on FFS height. The mechanisms responsible for the above behavioural modification are unknown, motivating the current thesis. The evolution of primary stationary crossflow instability, in its linear growth phase, is studied through a spanwise invariant, synthetic, and idealized rapid base flow deformation imposed by changing the near-wall pressure distribution of a clean swept flat plate (possessing a favourable pressure gradient) via Gaussian-like pressure variations. An energy balance framework is developed that identifies the production term's behaviour as the differentiator between regions of perturbation growth and decay. The behaviour of the production term is described by two competing mechanisms, the first controlled by wall-tangential base flow shear and the second controlled by wall-tangential base flow acceleration or deceleration. The balance of these mechanisms shows that perturbations grow faster than the clean case in regions of wall-tangential base flow deceleration and slower than the clean case in regions of wall-tangential base flow acceleration. Perturbations are even found to attenuate in some cases when a region of wall-tangential base flow acceleration follows a region of wall-tangential base flow deceleration. The modification of energy transfer mechanisms brings into question whether the initial modal stationary crossflow mode deviates from modal character on interacting with the base flow deformation. The Orr mechanism is shown to identify differences in perturbation behaviour from local modal character. However, criteria from the literature hint towards an absence of any non-modal effects. Finally, the extent to which a synthetic idealized base flow deformation mimics the effects of an FFS on deforming the base flow and changing trends of stationary crossflow instability evolution is tested to show the applicability of methods developed in the thesis to instances of natural base flow deformation. The progress in understanding mechanisms by which a deformed base flow affects the linear phase of primary stationary crossflow instability growth leads to suggestions on devices that can be tested to delay this phase of instability growth. These devices could potentially also delay subsequent stages of instability growth and hopefully lead to the development of novel transition delay techniques.

Recent years have seen an increase in studies focusing on data-driven techniques to enhance modelling approaches like the two-equation turbulence models of Reynolds-averaged Navier-Stokes (RANS). Different techniques have been implemented to improve the results from these simulations. In particular, the main focus has been on overcoming the limitations implied by the Boussinesq assumption. This has been approached by using machine learning techniques as a way of discovering new formulations that could overperform when compared to traditional models.
Despite promising results for steady RANS simulations, little has yet been investigated in URANS applications. In this dissertation, this lack of research will be addressed. The main ideas are then, first, to see how the available information in URANS simulations can be used to improve the anisotropic Reynolds stress tensor prediction, and second if and how this can be done by using a sparse regression technique, whose framework is known as SpaRTA. A procedure involving the triple decomposition of High-Fidelity velocity fields is applied, aiming at finding a model exclusively for the stochastic component of the anisotropy. The test case which is considered is the flow around a cylinder at Re=3900. The High-Fidelity data was collected by running Large Eddy Simulations in OpenFOAM, after which the velocity was split into its different components through Proper Orthogonal Decomposition.
A priori results have shown good performance of the trained models, outperforming the Boussinesq assumption both in the prediction of turbulence componentiality and also on the values of the single anisotropy components.

Experimental measurements are performed on a 45 degree swept flat plate model at the low speed laboratory (LSL) at the Delft University of Technology, in a low turbulence environment to stimulate the development of stationary crossflow. The swept flat plate model is equipped with two linear manual stages to create forward and backward facing steps. Preliminary measurements characterize the pressure gradient over the swept flat plate model under study. In the preliminary study , hot-wire anemometry (HWA) measurements characterize the flow over the swept plate without steps over a large chordwise domain with and without forcing by discrete roughness elements (DREs). The DREs are spaced at a spanwise wavelength corresponding to the overall maximum N factors from LST. The mean velocity contours and N factor trends presented in these measurements reinforced the need for DREs to control flow. Spectral content is monitored and the frequency bands associated with probe vibration and travelling crossflow interaction were delineated. Infrared thermography was employed to observe the movement of transition front with varying step heights and initial crossflow amplitudes. When the DRE height increases, the transition front moves upstream consistently for all step heights. Furthermore, when the DRE height is kept a constant , but the array is moved upstream and downstream of the neutral point, the transition front moves upstream for all step heights. In order to observe the flow in the vicinity of the step, HWA was once again used to quantify the interaction of crossflow with FFS. The clean, short FFS and supercritical step height configurations identified from the IR study, are studied for two initial amplitudes. For the supercritical step configuration, bandpass filtered fluctuations are found to align with a high wall normal and spanwise shear region which has been identified in previous work. It is postulated to be associated with a vortex shedding mechanism, for which frequency bands are delineated. Estimates of the range of recirculation bubble length were made and a flapping frequency range was also demarcated. In this study, a vortex shedding scenario is used to explain the presence of these near wall fluctuations. To conclude the report, recommendations are made for extending the present study for future work.

Following recent proof of swept wing laminar-to-turbulent boundary layer transition delay through base-flow modification using AC-DBD plasma actuators, an experimental investigation is performed to study the responsible physical mechanism. An experiment is conducted on a 45° swept wing in the anechoic wind tunnel at the Delft University of Technology. Stereoscopic PIV is employed to study the influence of the AC-DBD plasma actuator on the boundary layer velocity profiles. It is found that the AC-DBD plasma actuator is capable of directly decreasing the cross-flow component for various operating conditions. Moreover, as the Reynolds number is decreased, the authority of the actuator increases accordingly. In a second wind tunnel experiment in the same facility, the influence of the actuator on the cross-flow instability is monitored by performing 2D2C-PIV measurements at consecutive chord locations. As a result of the body force induced by the actuator, the cross-flow vortices shift position and feature a reduction in amplitude for the high frequency case of 5 kHz. Lower carrier frequencies are found to induce turbulent wedges due to the presence of strong localised plasma discharges. As a consequence, successful cross-flow instability control highly depends on the chosen carrier frequency and voltage, as well as the actuator lifetime and configuration. Although the changes in the transition front can not be observed here due to the consulted wing model, it is believed that the presented results provide an adequate overview of the influence of the AC-DBD plasma actuator on the cross-flow instability and add to the body of knowledge concerning AC-DBD plasma actuators and their application as flow control devices.