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. K-H type of 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 (450Hz ≤ 𝑓 ≤ 3000Hz), whilst the numerical high-frequency content (3500Hz ≤ 𝑓 ≤ 9000𝐻𝑧) 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 the 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.@en

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.

To achieve delay of laminar-to-turbulent transition of boundary layers, the application of acoustic metamaterial concepts is investigated. This is done analytically, numerically and experimentally by studying the interaction between a Helmholtz resonator and a TS wave. A Helmholtz resonator can serve as a meta-atom in a metamaterial. The desired wall controller property of a metamaterial to suppress TS waves is evaluated. This study shows that TS wave attenuation is achieved if the wall-normal velocity from a metamaterial is in-phase with the TS wave pressure. TS wave amplification is obtained if the pressure and velocity are more than (approximately) 90 degrees out of phase. The p-v phase relation corresponding to a Helmholtz resonator only remains in the amplification regime. Therefore, the conclusion is drawn that a single Helmholtz resonator can only amplify TS waves. If an acoustic metamaterial concept is found where pressure and wall-normal velocity are in-phase, this can be useful in boundary layer instability control.

This thesis concerns a numerical investigation for the stabilization (delay of laminar-to-turbulent transition) of ows over a at plate with a meta-unit of a possible metamaterial. The type of instability studied is that due to small disturbances in an incompressible, two-dimensional boundary layer. The linear growth of the disturbance in the form of a wave (also called the Tollmien-Schlichting wave) was of particular interest. There is a newly emerging class of structures called "metamaterials" which are capable of controlling the properties of classical waves (electromagnetic/optical/acoustic). This thesis investigates whether or not one of these materials are capable of attenuating TS waves. A literature review was conducted on three types of metamaterials, namely those based on Helmholtz-resonators, membranes structures and phononic crystals. The mechanisms for which the three aforementioned metamaterials attenuate acoustic waves are resonance, anti-resonance and Bragg's scattering respectively. Based on the likelihood of successfully achieving damping of TS waves, the phononic crystal was chosen to be further studied…

The location of laminar to turbulent transition is an important consideration as turbulent flow is associated with higher skin friction drag and, by extension, lower fuel economy. In unswept boundary layers, natural transition proceeds by the amplification of Tollmien–Schlichting waves. Tollmien–Schlichting waves are convective instabilities with spanwise oriented vorticity. The amplification of these instabilities/perturbations is sensitive to roughness elements, such as forward-facing steps. These surface imperfections are inevitable as steps, gaps, and humps are a byproduct of mismatch between panels of a wing. However, their interaction with Tollmien–Schlichting waves is not very well understood. Direct numerical simulation of the flow field around forward facing steps has been performed in this thesis to gain an in-depth understanding of the particular flow features that stabilise or destabilise the incoming Tollmien–Schlichting wave, with respect to a flat plate zero pressure gradient flow. The forward facing step is found to significantly distort the base flow, its effect scaling with the roughness Reynolds number in the upstream regime. This distortion of the base flow is observed to amplify the incoming instability, both upstream and far downstream. At the step location, however, stabilisation or destabilisation can be observed, depending upon the height of the step. The step causes the incoming Tollmien–Schlichting wave to split into two, just upstream of the step, and leads to two counter-rotating structures at the step location. The interaction of these structures influences downstream growth. Localised stabilisation is observed, at the step location, for step heights that are smaller than the boundary layer displacement thickness. Destabilisation is observed for larger step heights. The upstream base flow distortion is due to an adverse pressure gradient imposed by the forward facing step. The magnitude of the pressure gradient is found to scale with the roughness Reynolds number. The upstream amplification is due to the Tollmien–Schlichting wave encountering the distorted base flow. The response of the Tollmien–Schlichting wave to the distorted base flow is observed to scale with its wavelength. The ratio of the roughness Reynolds number to the wavelength ($\gls{Rehh}/\lambda$) is found to be the governing parameter for the upstream interaction of the step with the instability.

The laminar-turbulent transition of boundary layers is accompanied by an increase in friction drag. Laminar airfoils are a promising technology to reduce the emission of greenhouse gasses caused by the aerospace industry. The transition scenario in airliners is dominated by stationary crossflow vortices whose evolution is strongly nonlinear. This transition is preceded by the eigenmode growth of instabilities, rise of secondary instabilities and development of turbulent spots that eventually lead to a fully turbulent boundary layer. Exponential growth starts at infinitesimal amplitudes at which linear methods, e.g. the Orr-Sommerfeld or Linear Parabolized Stability Equations (LPSE), can be exploited to assess the growth of instabilities. As the energy of these instabilities increases, interactions become important and a nonlinear method capable of calculating wave-triad interactions should instead be used to correctly predict their evolution. This thesis concerns the development of a tool capable of solving the Nonlinear Parabolized Stability Equations (NPSE) and introduces a new method for the introduction of harmonics during the simulation that follow from nonlinear interactions. This method accounts for nonparallel and nonlinear effects in an Inhomogeneous LPSE (ILPSE) framework by modeling the interactions as a forcing term in this equation. The amplitude of the mode prior to its introduction is modeled via the nonlinear estimation of the growth rate. An investigation in mode introduction schemes proved that the introduction of modes severely affects the results from stability analyses. The ILPSE method reduces initial transients by correctly accounting for a harmonic’s history using a growth rate estimate from the other modes in the system. The improved introduction scheme severely increases the accuracy of the introduction amplitude and growth of higher harmonics.

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.

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.

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.