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.

A robust and flexible numerical framework was developed for modelling of dielectric-barrier discharge applicable to plasma actuated flow control. In OpenFOAM, a solver was built which solves convection-diffusion equations for ionized species coupled to Gauss’ law for electrostatics and includes a chemical model to account for the various chemical reactions taking place in non-equilibrium cold plasma. The framework relies on a predictor-corrector method and was verified and qualitatively validated in 1D using a volumetric cathode-fall glow discharge case. Quantitative validation in 1D and extension to 2D was recommended for future work as steps towards increased capability of the solver.

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.

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%.

In the context of the ongoing research regarding boundary layer transition, this project aims to study the effects of isolated cylindrical roughness elements on the transition from laminar to turbulent flow in a swept wing boundary layer. The project entails experimental campaigns: the 45 degrees swept wing model is placed inside the test section of the Low Turbulence Tunnel (LTT) and millimeter-sized roughness elements are attached to it, so that their interference with the boundary layer flow can be studied. The main goal is to investigate the flow features generating aft the roughness elements and their development into turbulent streaks. Moreover, this project aims to understand and quantify the effects of different roughness sizes on transition mechanisms, in the specific case of a three-dimensional boundary layer. This is achieved by means of different experimental techniques, such as PIV, HWA and thermographic flow visualization. The obtained results lead to a comprehensive description of this type of flows, with several phenomena changing their characteristics due to a change in parameters (freestream speed, roughness diameter and roughness height). Velocity field, spectral analysis of the signal, wedge width evolution, instability modes and velocity fluctuations in time were investigated with different parametric conditions, in order to understand how the flow was affected by them. Furthermore, the study allowed to evaluate the use of non-dimensional parameters such as roughness Reynolds number and aspect ratio for the prediction of the flow topology. The complexity of the phenomena involved is underlined in the conclusions of this research, with the interaction of several flow features making it a broad and interdisciplinary field of research.

The high efficiency of open rotor propulsion has led to a comeback of propeller propulsion systems in recent years. The main issue preventing wide application of the technology is its high noise level. One of the main contributors to this noise is the pressure fluctuations generated during the interaction between a propeller tip vortex and a downstream surface. The goal of this research project is to create a better understanding of the vortex- surface interaction, which would allow for more effective noise mitigation strategies to be developed. An experimental approach was adopted to quantify the effect of a number of governing parameters on the pressure fluctuations over the surface induced by the vortex. Furthermore, an attempt was made to develop a conceptual model describing the vortex-surface interaction in order to study the relative contributions of the sub-phenomena of the interaction. The accuracy of the conceptual model was assessed using the data obtained from the experimental campaign. From the results it was clear that the low pressure vortex core was the primary mechanism that generated fluctuations over the surface. The vortex path was clearly visible and dominant over a large part of the chord. In the slipstream region the wake generated the largest pressure fluctuations. The effect of the wake was contained at the leading edge, up to 4% of the wing chord. Large discrepancies between the unsteady pressure, both in magnitude and in spatial distribution, were observed. Three contributing factors to the discrepancies were identified, namely the overestimation of the induced angle of attack effect, the use of the Lamb-Oseen vortex model and the method of determining the vortex core radius. These discrepancies implied that the conceptual model could not be used for its intended purposes. The results of this study show that the advance ratio and incidence angle are the most critical parameters governing the vortex-surface interaction. Although the former is difficult to incorporate into noise mitigation strategies, the latter can be used as a parameter in the initial design phase. Interestingly, the geometry of the airfoil seems to have limited effect on both the magnitude and distribution of the interaction. Although the conceptual model discussed in this report is not accurate, the development of such a simplified model would still contribute a lot to the body of knowledge. It could be used to evaluate the relative contributions of the different vortex-surface interaction sub-phenomena, or to perform a 'parameter sweep' to investigate the effect of the governing parameters on the unsteady pressure distribution and magnitude.

An aneurysm refers to the local dilation or ballooning of a blood vessel and if left unchecked, almost every aneurysm continues to grow in size until it eventually ruptures. The abdominal aorta is the most common location for an aneurysm to form and an abdominal aortic aneurysm (AAA) can turn lethal in the event of rupture. Since the current clinical practice to classify AAAs is on the basis of their size (diameter), it was decided to investigate the impact of varying geometric parameters viz., the expansion ratio (ER) and the expansion angle (EA), on the pulsatile flow field in a hypothetical axisymmetric AAA (4A) geometry and in particular, investigate the role turbulence. All the combinations of ER and EA lead to a total of 8 cases. The simulations show that a vortex ring is shed just after peak systole in every case. The vortex ring, at inception, is similar for all cases within visual limits. As the ring travels downstream, an azimuthal instability sets in that grows with time. This is the short wave or the elliptical instability usually associated with a vortex pair perturbed by an infinitesimally small amplitude sinusoidal wave. The number of waves formed along the circumference of the ring as a result of the instability varies from case to case but in general, the number of waves decreases as the ER increases. The growth of this instability, along with interaction of the ring with the remnants of the flow field form the previous cardiac cycle, cause it to break down into smaller vortices thereby generating turbulent fluctuations. Further, subtle differences between various cases regarding the vortex ring breakdown mechanism are observed and discussed. Investigations are also carried out to identify recirculating or dead flow regions as in the context of an AAA, such a region indicates the presence an intraluminal thrombus (ILT) which is a commonly found feature in AAAs. An ILT deprives the surrounding arterial tissue of oxygen thereby weakening the aortic wall and increasing the risk of rupture. From the simulations, it could be seen that the inflection regions of the 4A geometry are most prone to the formation of an ILT which is in agreement with the physically observed locations. Finally, the oscillatory shear index (OSI) is investigated to see the influence of turbulence on a hemodynamic parameter. It was found that for the geometries tested, there is no clear co-relation between the two. From this thesis, numerous conclusions can be drawn. But perhaps the most important conclusion is that the flow field in an AAA is very complex and sensitive to various input parameters such as the ER and EA. Although flow parameters such as the Reynolds and the Womersley number are kept constant across all the cases, existing literature clearly highlights the dependence of the flow field on them. As such, the desire to be able to formulate general co-relation trends between various flow field quantities in AAAs is wishful thinking at best.

One of the main goals of laminar flow control is to reduce skin-friction drag by delaying the onset of laminar to turbulent transition. Over unswept wings, the leading cause of this is the growth of flow instabilities called Tollmien-Schlichting(TS) waves. This thesis aims to test the feasibility and performance of a designed Linear Quadratic Gaussian (LQG) controller in attenuating TS waves, in an experimental setting. A modified flat-plate test setup with an adverse pressure gradient was used to generate and characterize these waves, while control was performed using a Dielectric-Barrier Discharge (DBD) plasma actuator. Embedded microphones were used to measure the pressure fluctuations within the boundary layer, while Particle Image Velocimetery (PIV) was used to quantify the flow. The performance of the controller was investigated at its nominal and off-design conditions (robustness), as well as its comparison to open-loop continuous forcing. For the nominal design conditions, a reduction in the spectral energy of the peak frequencies of four orders of magnitude are observable for the closed-loop case, two orders more than the open-loop case. The root-mean-square (RMS) of the downstream signals showed an overall maximum additional reduction of 55% of the pressure fluctuations, compared to the open-loop case. The controller was more robust at lower peak-to-peak voltages, with an overall 30-60% reduction in RMS of the pressure fluctuations, compared to open-loop forcing .This study demonstrates the feasibility of using LQG model-based techniques as a viable flow-control strategy in damping flow-instabilities, and is an option worth further investigation.

Big wind farms with big wind turbines are more cost effective and produce greener electricity compared to smaller ones. This is one of the reasons for the growth of wind turbine size during the last 3 decades. However, as wind turbines become bigger, their blades become longer, thicker and heavier. This results in larger unsteady loads on blades which is an important limitation for their life time and their size growth. Flow control has emerged as the promising solution both for improving the aerodynamic efficiency and controlling the unsteady loads on wind turbine blades. DBD plasma actuator is an active flow control mechanism that has shown high potentials for unsteady load control with the capability to change lift coefficient to a significant amount with a very fast response time. The current research first intends to identify the variations of angle of attack and lift coefficient on wind turbine blades as a result of gravitational loads, mass and aerodynamic imbalances, turbulence, wind shear, yawed inflow and tower shadow and investigate their corresponding frequencies and the fatigue damage from the blade root bending moments. Then, (single and multi) DBD plasma actuator with different configurations will be used on three different trailing-edge shapes (round, half-round and sharp) of modified version of ’NACA64-2-A015’ airfoil to control the aerodynamic loads via circulation control. This is managed either by manipulation of Kutta condition or acting as a virtual Gurney flap. Furthermore, it is intended to investigate the correlation between the frequency of actuation, frequency of vortex shedding and the amount of lift enhancement.

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.

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 (Reτ=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 Cμ=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 (Cμ=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.

Research into constructing reduced-order models (ROM) to reduce computational cost or to interpret complex datasets from numerical or experimental sources has steadily been gaining popularity over the past few years. Many methods currently exist to construct ROMs. The proper orthogonal decomposition (POD) is one of the most widely adopted methods used for these purposes in the field of fluid dynamics. The POD modes favor high-energy structures due to the formulation of the POD however. This may result in an inaccurate ROM, as it might be possible that dynamically relevant, yet low-energy structures are truncated in favor of higher-energy structures. In recent years, an alternative goal-oriented method has been proposed. This goal-oriented method not only has less limits on the goal function that is optimally represented relative to thePOD, the resulting reconstruction of the flow field is model constrained as well. Both these properties are conjectured to lead to more dynamically relevant modes. This goal-oriented, reduced-order modeling (GOROM) technique has been formalized for the advection, Burger’s and incompressible Navier-Stokes equations in 1D and the Stokes equations in 2D, all with homogeneous boundary conditions. This work extends the formulation to the 3D incompressible Navier-Stokes equations with inhomogeneous boundary conditions and provides a parallel implementation. The method is applied to a DNS dataset of a 2D transitional boundary layer with and without a forward-facing step present and compared to the performance of the POD. It is found that the GOROM modes provide a basis which is numerically more accurate than the POD modes for most cases, yet do not provide additional insight into the flow physics. The results presented nevertheless show promise for further applications to construct more accurate and goal-oriented ROMs with the GOROM technique, especially for problems with nonlinear interactions and nonlinear goal functions.

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.

This project proposes a series of experiments that involve plasma synthetic jet actuators. The first experiment will perform a jet characterisation experiment that will research the effect of orifice geometry on the overall performance of the actuator. The second experiment will built upon the first experiment and will use a plasma synthetic jet array to combat leading edge separation of a NACA 0015 airfoil at Re=1.7⋅105 and U∞=10 m/s and improve the overall performance of this particular airfoil. Special focus will be put in uncovering the underlying mechanics of plasma synthetic jet actuation operating in leading edge separation conditions and how performance is dependent on the actuation frequency. From the jet characterisation experiments a clear performance trend of actuator efficiency with respect to the converging cone angle (θ) is found. The optimal optimal orifice angle is expected to lie between 45º<θ<69º. These geometries experience ∼20% higher jet velocities than the baseline 'straight' orifice, which results in larger mass expulsions and an overall more efficient operation of the actuator. Additionally it is found that adding a small diverging section to the orifice improves upon the electro-mechanical efficiency of the plasma synthetic jet actuator. PIV measurements show that this is due to the increased effective orifice area which allows for higher mass flows through the orifice but also the jet velocities remained of similar order as the optimal converging geometries. The flow control experiments show that plasma synthetic jet actuators can indeed improve the performance of a NACA 0015 airfoil at Re=1.7⋅105 and U∞=10 m/s. The force balance measurements show that PSJ actuation suppresses the hysteresis loop present when actuation is absent. Furthermore the angle at which maximum lift is achieved is shifted by ∼7º increasing the maximum achieved lift by ∼23%. Additionally, flow separation can be delayed by about 2º reducing the drag by about ∼40%. Furthermore, the PIV measurements show the mechanisms behind flow separation control. At moderate stall angles flow reattachment is feasible if the actuation frequency is high enough. At higher angles of attack the separation point moves upstream of the actuators and renders the array incapable to suppress flow separation. However, at these conditions the actuators are still able to influence the separation region and higher frequencies, with an optimum of F*=1, are capable to suppress the separation area more. If the above-mentioned experiments translate to aeronautical applications plasma synthetic jets might be a game changer when it comes to demanding flight conditions. Not only is plasma synthetic jet actuation capable of diminishing the hysteresis effect it is also capable of considerably increasing the lift and decreasing the drag forces. These effects can considerably improve the safety of aircraft as the omission of hysteresis can reduce unwanted unsteady loads that advance structural fatigue and the higher lift coefficients reduce the need of high lift devices allowing them to become smaller and less complex in the future. Overall this allows aircraft to fly at more demanding flight conditions than previously feasible.

Hybrid laminar flow control (HLFC) reduces skin friction drag, by combining boundary layer suction with pressure gradient tailoring to delay boundary layer transition. This research addresses two key developments required to perform the aerodynamic design of suction-type HLFC components. First, an existing numerical transition prediction tool was made compatible with boundary layer suction, by incorporating suction in its boundary layer solver and by ensuring that an appropriate grid of disturbance frequencies is evaluated with linear stability theory. Second, the scalability of the pressure losses associated to the perpendicular flow through large-scale and actual-scale perforated sheets was investigated experimentally. This development is required to predict the pressure losses of generic HLFC surfaces. It was found that the scalability of these pressure losses is limited, because the frictional losses inside holes decrease continually with decreasing hole diameter. In contrast, the inertial losses across holes reach a steady value with sufficiently many holes.

The research presented in this thesis focuses on boundary layer separation and, more specifically, on the related physical mechanisms, such as laminar to turbulent transition, with a vision towards effective separation control. The work is divided in three parts, each corresponding to a different separation scenario. The supporting data is experimental, obtained by means of particle image velocimetry. The first and major part regards the phenomenon of laminar separation bubbles. The spatial and temporal response characteristics of a flat plate laminar separation bubble to impulsive forcing are first investigated in order to shed light on the processes of flapping and bursting. The impulsive disturbance is introduced twodimensionally with a dielectric barrier discharge plasma actuator. The disturbance develops into a wave packet that causes rapid shrinkage of the bubble in both upstream and downstream directions. This is followed by bubble bursting, during which the bubble elongates significantly, while vortex shedding strength in the aft portion of the bubble is reduced. Duration of the recovery of the bubble to its unperturbed state is independent of the forcing amplitude. At the same time, linear stability analysis shows that the growth rate and the frequency of the most unstable mode decreases for increasing forcing amplitude. Throughout recovery, growth rates are directly proportional to the shape factor, indicating that bursting and flapping mechanisms are driven by altered stability characteristics due to variations in incoming disturbances. It is found that the stability of the flow changes only when disturbances interact with the shear layer breakdown and reattachment processes, supporting the notion of a closed feedback loop.@en

The research presented in this booklet focusses on the cross-flow instability. Applying traditional and advanced flow diagnostics, the boundary layer evolution is studied in detail. The topology and evolution of both primary and secondary instability mechanisms is revealed with unprecedented detail for experimental research paving the way for new advanced-diagnostics investigations. Important confirmations of the outcomes of past experimental, numerical and theoretical studies are achieved together with the description of a newly-reported flow phenomenon. The latter consists of a low frequency motion of the "stationary" primary vortices. While this phenomenon is considered not relevant for the transition evolution, it is deemed important for experimental investigations as it encompasses very high levels of turbulent kinetic energy. Advanced flow control experiments based on alternating current dielectric barrier discharge plasma actuators are also performed following different instability control approaches. The primary instability is conditioned by the external forcing either in the wavenumber spectrum (by inducing selected spanwise modes) or in intensity (by weakening or enhancing the cross-flow velocity). The secondary instability modes are conditioned in the frequency spectrum and phase. These efforts achieved the intended scopes. Although, when selected stationary modes were forced, the boundary layer fluctuations were enhanced. These fluctuations can directly cause the turbulent breakdown vanishing the beneficial effect of the performed instability control. The cross-flow forcing, making use of newer actuators reaching higher frequencies, resulted successful yielding transition promotion or delay depending on the forcing direction. @en

The market growth expected for commercial aviation in the coming decades and the increasing social awareness regarding the effects of global warming are driving significant technological developments necessary for emission reduction in future transport aircraft. From the aerodynamics perspective, a significant increase in aircraft efficiency can be obtained by applying Laminar Flow Control (LFC) techniques. The objective of LFC techniques is to reduce the skin-friction drag component by delaying the laminar-turbulent transition through the stabilisation of boundary-layer instabilities. Relevant to high-subsonic transport aircraft is the development of Crossflow (CF) instability, which manifests as a series of co-rotating vortices inside the boundarylayer flow on swept aerodynamic surfaces...@en

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.@en