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 recent years, research for turbulent boundary layer (TBL) has shifted towards developing novel approaches to reduce drag. This shift is driven by new industrial standards emphasizing fuel-efficient and environmentally friendly air transportation. As wall-bounded turbulence plays a significant role in the overall drag experienced by an aircraft, any understanding and control authority on associated flow structures becomes vital for more optimized and efficient flow control techniques to reduce drag. In this context, conventional passive flow control techniques have been widely used as they are simple and do not require any external power supply. However, active flow control techniques have demonstrated the potential for greater effectiveness, which offers more substantial drag reduction. Existing active control methods, such as wall oscillation, traveling wave, blowing, and suction, have been previously studied to reduce the viscous drag in the TBL. However, the efficacy of such control strategies tuned to outer layer large-scale motions (LSMs) at higher Reynolds number (Re) flows has emerged as a promising avenue for achieving significant drag reduction. In contrast, control techniques targeting the inner layer small-scales encounter limitations due to restricted control authority and the difficulties associated with miniaturizing hardware at higher Re flows.

LSMs convecting in the logarithmic region (log region) of TBL play a critical role in the dynamics of wall-bounded turbulence, as they carry a significant portion of the turbulent kinetic energy (TKE) [Abbassi et al., 2017]. The active wall oscillation technique tuned to the frequency of the LSMs has proven to be energy-efficient, as drag reduction increases with the frictional Reynolds number (Reτ ) and the contribution of LSMs to the TKE and wall shear stress increases with the Reτ [Marusic et al., 2021]. In this thesis work, the receptivity of LSMs to active large-scale control strategy is assessed via a spanwise array of wall-normal jets in a TBL. Multiple wall-normal jets manipulate the flow over a flat plate by introducing a spanwise traveling wave, which aims to mimic wall oscillation tuned to the large scales in the log region of the TBL.

To assess the efficacy of the suggested control strategy with spatial and temporal tuning, particle image velocimetry (PIV) and hot-wire anemometry (HWA) are employed to quantify second-order turbulence statistics and analyze the organization of the coherent patterns introduced in the flow via two-point correlations. In addition, the effect of different control cases derived from different tuning and control strategies is analyzed at Reτ = 2227. The TKE production plots show that control cases targeting the LSMs effectively attenuate the TKE production near the wall. Furthermore, based on the analysis of the spectral energy plot of streamwise velocity fluctuations, it is observed that the energy content associated with larger length scales convecting in the middle of the log region decreases in the controlled case compared to the uncontrolled base case. This finding suggests that certain control cases notably impact the organization of LSM and second-order turbulence statistics far downstream of the actuation point. Additionally, the reduction in TKE production near the wall represents a reduction in wall shear stress and viscous drag. However, additional research must be conducted in the future based on the current findings to reach a definitive conclusion regarding the efficacy of individual control cases.

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

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

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

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

Turbulent drag reduction (DR) is a crucial research area, as this can contribute to significant energy and emission reductions in various industrial and transportation applications. One such technique is the active flow control method of spanwise forcing. Spanwise forcing introduces a spatio-temporal spanwise wall oscillation as a boundary contrition. Consequently, a phase-varying spanwise velocity profile is generated, that poses useful interaction with the near-wall turbulence resulting in a drag-reduced flow state. DR values of over 40% have been found in both numerical and experimental realisations.

The objective of this thesis is twofold. Firstly, the idea is proposed that spanwise forcing can be realised passively by geometric surface modifications to create a spatially varying spanwise velocity profile. Secondly, supporting the first objective is the idea of investigating active flow control, in this case spatial spanwise forcing, which can aid in the development of passive techniques. The underlying philosophy is gaining understanding of the key factors and mechanisms that contribute to DR by studying active techniques. Subsequently, these leanings can be applied to advance the development of passive techniques. This research rationale is combined with a new pathway for turbulent DR. By actuating the outer-scaled low-frequency turbulent structures in the logarithmic layer, instead of the conventional approach targeted at high-frequency inner-scaled structures. Unlike the decreasing trend observed in inner-scaled actuation (ISA), outer-scaled actuation (OSA) exhibits a positive correlation between DR and Re_τ, making it particularly useful for practical applications that operate under high Reynolds number conditions, such as long-distance pipelines or aviation.

The first part of the thesis investigates the DR and OSA potential of passive techniques. Three techniques are considered: dimples, sinusoidal undulations and oblique wavy walls. Eight test plates are realised for direct force measurements and a detailed investigation into the scale-specific streamwise turbulence kinetic energy using hot-wire anemometry (HWA). From the direct force measurements, the DR trends with Re_τ are inconclusive due to uncertainties originating from a correction term and the results falling within the 95% confidence interval. A slight reduction of outer-scaled energy is observed for the sinusoidal undulations, hinting at effective OSA. Furthermore, one of the oblique wavy walls shows a broadband energy reduction in the wall-normal region 10 ≤ y⁺ ≤ 100.

The second part of the thesis is focused on active flow control, specifically, the method of spatial spanwise forcing. A research gap was identified for spatial forcing, with no experimental realisations to date and only a number of numerical studies. Furthermore spatial forcing is of interest since it shows the closest analogy to a passive counterpart. An experimental setup is proposed, titled the steady spanwise excitation setup (SSES). The experimental setup realises spatial forcing using a series of spanwise running belts that run in alternating spanwise directions so as to generate a spatial square wave forcing. The thesis presents a proof of concept of the aerodynamic working using a prototype setup with four belts. Stereo particle image velocimetry (PIV) measurements showt hat a significant flow control effect can be realised with this type of forcing, with a maximum DR value over 39%. Accordingly, the turbulent stresses are significantly reduced. The spanwise velocity profiles are in qualitative agreement with the model based on the laminar solution, with an almost exact match at y⁺ ≤10.

Flat surfaces with arrays of wall-normal pores called micro-cavity arrays have shown potential to reduce turbulent skin friction drag. Their designs have been inspired by acoustic liners which are sandwich panels consisting of a honeycomb core, a perforated top plate and a solid back sheet used on aircraft engine nacelles for noise abatement. Research on acoustic liners has paved the way to identify the parameters that are crucial for optimising the design of micro-cavity arrays. Spanwise rectangular grooves, a special case of cavities, have also shown potential to reduce skin friction drag. This thesis aims to validate the reported drag performance of micro-cavity arrays and grooves through experimental methods used in literature and direct force measurements. 

In this research, experimental results show that although pores and grooves are capable of reducing turbulence energy in the boundary layer, there is a significant drag increase with respect to a smooth reference. Based on measured drag performance and past research on cavity flows, it has been argued that the observed drag increase is caused by pressure forces acting on the inner walls of the pores and grooves. As past research on micro-cavity arrays did not perform direct force measurements, the contribution of pressure drag had not been investigated. Pressure drag is expected to be present and perhaps even dominate over skin friction drag reductions. Further research into the flow mechanics inside the pores is required to validate this argument and find a way to minimise the pressure forces. Wall-normal pores could still be an effective turbulent drag reduction technique if arrays are optimised for skin friction drag reductions that surpass pressure drag.

The flared folding wingtip (FFWT) is a concept presented by Airbus to increase the aerodynamic efficiency and provide a means of load alleviation. A gate-to-gate demonstration of this concept has already been presented with the AlbatrossONE, a scaled aircraft featuring the FFWT concept. In this demonstration, the aircraft fold the wingtips during taxiing to fulfil gate requirements and extends it before take-off for improved efficiency. In addition, they can be released during flight to provide load alleviation and reduce the roll rate reduction caused by the increase in the wingspan. The objective of this thesis is to further study the FFWT concept.

An aeroelastic wind tunnel experiment to identify the influence of the wing stiffness and hinge release threshold on the gust load alleviation performance of a folding wingtip design is presented in this study. Five models with different stiffness and tailoring properties are tested and the wing root bending moment at different conditions is compared to the response with locked hinge conditions to assess the impact on the gust load alleviation capabilities of the folding wingtip.

The results show that the structural properties do not have an important impact on the peak load alleviation but the hinge release threshold and timing do. Releasing with the correct timing can reduce significantly the peak loads. However, the dynamics of the system are affected by this release: the flutter speed is decreased and, although the performance can improve, load oscillations increase, which can be considered detrimental for reasons such as fatigue.

Dimples are shallow surface indentations that have been recently considered as a passive viscous drag reducing technique for turbulent boundary layers. Studies so far were limited to incompressible low-Re flows and display little consensus on whether dimples can actually produce net improvements in total drag. Conversely, dimple geometry can be regarded as the inverse of transonic bumps, a matured wave drag reduction. In light of this, this thesis sets off to analyse the flow structures that arise over dimples at transonic speeds provide, serving as an initial assessment of their wave drag reducing potential.

A transonic flow was reproduced experimentally using an asymmetric nozzle. Then, the performance of five dimple designs was evaluated against the criteria adopted for transonic bumps. Measurements included Schlieren, surface pressure and PIV. Results reveal that small spherical dents seem to produce spanwise excitations that subdue the detrimental effects of expansion fans formed at their edges, and ultimately allow for a higher momentum retention across the interaction.

The development of air-breathing propulsion systems as a mean to propel high-altitude high-speed transport and single-stage-to-orbit (SSTO) vehicles promises to be a viable alternative to chemical rocket propulsion. In order to enhance mixing, an oblique shock wave is introduced into the combustion chamber as a source of baroclinic vorticity as well as a static temperature and pressure augmenter. An experimental campaign was conducted in the ST-15 supersonic test facility at Mach 2 to analyse the effect of three main control variables: the jet momentum flux ratio, the flow deflection angle and the impingement position of the shock on the jet plume. Measurements were acquired with Schlieren photography and stereo PIV techniques. Results suggest that, while near-field momentum-driven mixing remains unaltered following the introduction of the impinging shock wave, mid-to-far-field mixing mechanisms do change. An increase in the jet plume elevation was observed together with one in lateral expansion as a consequence of the introduction of a shock wave. Also, the formation of a strong shear layer downstream of the jet was observed, which acts as a source of vorticity to promote entrainment towards the jet mid-field. A stronger wave was noticed to produce more optimistic results for the mixing performance. This effect was seen to decrease with the introduction of a weaker shock or by shifting the strong shock downstream.

Turbulent drag is the largest source of fuel consumption in aviation and forms a significant contribution to climate warming. Recent numerical studies have suggested the ability of surfaces with streamwise-preferential permeability to reduce turbulent friction, and a theoretical framework behind the working mechanism has been proposed. This work is the first experimental study on this concept in air and in which the permeability requirements for predicted drag reductions are met. Three different physical realisations of streamwise-preferential permeable substrates were manufactured: two seal furs, a substrate with unidirectional fibres, and a 3D-printed structure. They were investigated through wind tunnel experiments using direct force measurements and planar (2D-2C) particle image velocimetry (PIV). Results show an increase in drag for all test specimens and suggest that drag sources other than surface friction are non-negligible. One-point turbulent statistics show an overall increase in turbulence, mainly wall-normal velocity fluctuations enabled by the wall-normal permeability, which results in higher Reynolds stresses. No significant flow modulation in terms of turbulent events or change in coherent structures is observed. The measurements and results are a first of its kind and none of the theoretical framework predictions on drag reduction or flow modulation are found. Based on analytical derivations relating characteristic length scales to permeability, the validity of the framework and its assumptions in experimental settings is questioned. It is thought that there is an inherent mismatch between the pore size assumption and the virtual origin approach taken in the framework. All things considered, it is deemed unlikely that turbulent drag reduction by means of streamwise-preferential permeable surfaces is feasible in experimental settings. Nevertheless, given the relatively small increase in drag for the 3D-printed specimen, streamwise-preferential permeable surfaces might be interesting for other (flow control) purposes involving turbulent boundary layers.

In experimental aerodynamics, pressure is an important parameter which provides insight into the various features of the flow around the test object. By obtaining the surface pressure information, various inferences can be drawn such as flow separation, local flow velocity etc. Traditionally, the most common method of obtaining pressure has been through direct measurements by incorporating the test model with pressure taps, which makes the model expensive and complicated. However, often there arises a requirement to obtain qualitative information and visualise the aerodynamic features during the test. Particle Image Velocimetry (PIV) solves provides this qualitative information as well as quantitative information in the form of a flow field. Pressure can be further computed using this velocity field. This technique holds a number of advantages over traditional wind tunnel measurements, especially with reduction in resources for manufacturing, ease of handling and non-intrusiveness. However, the conventional PIV technique came with the disadvantage of only having 2D2C information. Other techniques like stereo PIV and tomographic PIV were developed to tackle these limitations and the most recent development was that of Coaxial Volumetric Velocimetry (CVV). Along with the advent of Helium Filled Soap Bubbles (HFSB), this novel technique allows for large scale 3D3C measurements. This research is conducted to assess the feasibility of this promising technique to obtain surface pressure.
The flow field around an inverted wing in ground effect is investigated using CVV by robotic manipulation. Using HFSB as tracer particles, the experiment was carried out in the W-Tunnel of TU Delft. The wing was tested at two angles of attack, α = 5° and α = 8° and three ground clearance configurations
(h/c= 1, 0.6 and 0.3 at α = 5°) at airspeed of 10 m/s. The wing is equipped with a total of 30 surface pressure taps along the chord-wise direction, both on the pressure and suction side. The velocity field is obtained using Lagrangian particle tracking algorithm, Shake-the-Box and time-averaged velocity field is obtained by ensemble averaging of the obtained particle tracks. Static pressure is calculated using Poisson’s equation using the information obtained from velocity field. The pressure thus obtained is then compared with the readings obtained from pressure taps.
Such an assessment not only shows the feasibility of the technique but also highlights the shortcomings and potential improvements. To the best of the author’s knowledge through a literature study, this particular technique has not been used to obtain surface pressure information. Most of the work published in the field of CVV is to study flow features, mostly in the wake. Hence, it is expected that this research and the recommendations help to further development of such a technique in the future.

Transonic buffet is an aerodynamic instability that appears on aerofoils and wings at certain combinations of Mach number and angle of attack, and which limits the aircraft flight envelope. This thesis project presents an experimental
investigation of buffet employing stereoscopic PIV and Background Oriented Schlieren. By using four aerodynamic models, a 2D aerofoil and three 3D wings with sweep angles of 0^o, 15^o and 30^o, the differences between the 2D buffet typical of aerofoils and the 3D buffet typical of swept wings is studied, scrutinising the influence of the sweep angle. The results confirm that for the same aerodynamic conditions the 2D buffet experiences larger shock oscillations than the 3D case. A frequency analysis reveals a clear peak at 160 Hz for the 2D case, which is also visible in the 3D case for low sweep angles. As the sweep angle increases, this peak is replaced by a bump at higher frequencies.

Reduction of skin-friction drag over a fully developed canonical zero-pressure gradient turbulent boundary layer (ZPGTBL) subjected to spanwise oscillation is measured using planar particle image velocimetry (PIV). The experiments are conducted at Re_θ of 1000 and 1800, the chosen range of spanwise oscillations amplitude and frequency are around the optimum reported in literature studies (T⁺_osc = [100-700], A⁺_osc= [50, 150]). A high-resolution planar PIV apparatus is employed to measure the drag reduction directly by wall shear stress estimates on the oscillating wall. The measurement uncertainty of the drag estimates from PIV measurements is examined. The results show drag reduction in the order of 15% after 6 boundary layer thicknesses from the beginning of the oscillating section. Variations of the drag reduction follow the trends reported in literature. The PIV measurements enable the analysis in terms of Reynolds shear stresses, turbulence production and allow visualising vortex packets by vorticity. A pronounced drop of turbulence production in the range y⁺ = [5, 30] is observed. The vorticity analysis indicates a distortion from the well-known hairpin-packet arrangement, suggesting that the mechanism of hairpin auto-generation may be inhibited by spanwise wall oscillations

Flapping wings display complex flows which can be used to generate large lift forces. Flexibility in wings is widely used by natural flyers to increase the aerodynamic performance. The influence of wing flexibility on the flow can be computed using numerical analysis with Fluid Structure Interaction (FSI). However, there is a lack of open-source FSI methods. Therefore, a new implementation is developed for the well-known CFD code OpenFOAM to facilitate FSI with the multi-physics coupling library preCICE. The OpenFOAM adapter does not require changes to the OpenFOAM source code and is compatible with a variety of OpenFOAM versions. Structural modelling is performed using the open-source code CalculiX.

The method is validated using the laminar incompressible cylinder-with-a-flap benchmark for one steady and two unsteady cases. Good coherence is seen for the deflection and force generation, but the coupled method is sensitive to the eigenmodes of the structural model.

The influence of inertial, elastic and aerodynamic forces is quantified using a 2D wing. A sinusoidal flapping motion is imposed on the leading edge of the vertical wing. The inertial force on the wing dominates for high mass ratios and the wing deflection is rather independent of the flow. For a low mass ratio, the wing deformation scales with the increasing elasticity. The maximum lift and lowest drag were found for the wing with large flexibility and low mass so the passive deformation by aerodynamic forces creates a favourable shape for lift production.

Flexible translating and revolving wings at an angle of attack of 45deg show that chordwise flexibility decreases both lift and drag, however the lift over drag ratio is increased. The flow around both wings forms a coherent structure with a Root Vortex (RV), Tip Vortex (TV), Leading Edge Vortex (LEV) and Trailing Edge Vortex (TEV). The LEV on the revolving wing is stable for approximately up to half the span because vorticity is transported outward in the vortex core. The flowfield and LEV breakdown are consistent with experimental data of the same wing. The translating wing builds up circulation but the LEV detaches quickly near the centre of the wing. Chordwise bending reduces the angle of attack which decreases the distance to the core of the shed LEVs.

Recent decades, an ongoing trend has emerged in the upscaling of wind turbines in order to compete with traditional energy resources. For conventional wind turbines, aside from improving safety, the driving factors for new designs have always been increased efficiency and performance. This has led to novel wind turbine designs comprising thin-walled structures and light-weight composite materials that produce more power per unit. The current cost of wind energy is strongly dominated by operational and maintenance costs throughout the full lifetime of a wind turbine. Most wind turbines do not reach their design lifetime due to various reasons, from which fatigue failure is the most prominent one. The events responsible for gearbox or blade failure are caused by complex interactions between aerodynamics and structural responses that are inherently of unsteady nature. Aeroelasticity has become increasingly important for a safe and cost-effective design. Additionally, longer and lighter wind turbine blades undergo extremely large, low-strain deformations. A more accurate understanding of the aeroelastic behavior demands nonlinear analysis methods.

Most aeroelastic solvers in aerospace industry rely on linear structural models. This thesis work has modified the existing semi-nonlinear aeroelastic analysis method of the in-house developed CFD code at NLR by using Nastran's nonlinear structural module. Both analysis methods were utilized to obtain a converged static aeroelastic solution for two different operational conditions of the design curve. Subsequently, mode shapes of the deformed state of the structure were determined, to be used in the flutter analysis.

A 108-meters theoretically designed blade, provided by the Dutch blade design company We4Ce, was analysed using both methods. The results of static aeroelastic analysis and flutter analysis were compared to assess the effect of structural nonlinearities on the aeroelastic behavior. It is concluded that linear analysis overestimates the structural deformations and that pre-stressing due to nonlinear deformations and follower-forces alter the dynamic properties of the wind turbine blade. For the test case considered, the inclusion of geometrical nonlinearities resulted in a change of behavior of various modes. Naturally-low damped modes were affected negatively, eventually leading to dynamically diverging behavior. This thesis has successfully provided the first steps in the implementation of a nonlinear aeroelastic analysis for extreme scale wind turbines, including the capability of performing stability analysis.

The boundary layer transition location is a crucial design parameter in aerodynamics. The computational prediction of boundary layer transition in engineering applications is typically based on empirical models. These models require experimental measurements for calibration and validation purposes. For unsteady aerodynamic processes, the range of suitable boundary layer transition measurement techniques is traditionally limited to fast-response discrete sensor techniques, e.g. hot-film anemometry. Introduced in 2014, differential infrared thermography is an alternative approach to measuring unsteady boundary layer transition using thermal images acquired with an infrared camera. The application of this optical measurement technique reduces the experimental effort, but problems emerge when the temperature response of the surface under investigation is slower than the aerodynamic unsteadiness. The idea of differential infrared thermography is to evade this problem by subtracting subsequent thermographs and accrediting the largest difference to the moving boundary layer transition location. The working principle of the technique has been demonstrated in various experimental setups. In a heat transfer simulation, it has been shown that the image separation time between the subtracted thermographs should be as small as possible to avoid systematic measurement errors. An experimental validation of these results is performed in the present study, conducted at the DLR in Göttingen. In this study, the infrared radiation from a pitching airfoil model suction surface in the wind tunnel at Ma = 0.15 and Re = 1 ⋅ 10^6 is measured with an infrared camera. The pitching frequencies f and amplitudes a_1 are varied in a range from f = 0.25 Hz to f = 8 Hz and a_1 = 1° to a_1 = 8° to study their effects on the boundary layer transition measurements. The differential infrared thermography technique is applied with several different image separation time steps to confirm the findings of the heat transfer simulation. The time step size is optimized as a compromise between the measurement error and the signal strength. In the next step, two alternative boundary layer transition measurement schemes using infrared thermography are developed, based on a quasi-steady model and a heat transfer simulation result. The first newly developed approach involves the automated selection of the appropriate image separation time step for each data point over the motion period when applying differential infrared thermography. This adaptive approach outperforms the conventional fixed time step approach when applied to experimental data from the wind tunnel test. The second newly developed approach is termed local infrared thermography. When analyzing the thermographic measurement results at fixed model locations, the extraction of the extrema of the measured radiation signal yields instants of the motion period that correlate with the occurrence of boundary layer transition. It is demonstrated that the local infrared thermography approach can be extended to measuring two-dimensional boundary layer transition fronts. The analysis of these results provides insight into the origins of the differences between the results of differential infrared thermography and the reference measurements of the unsteady boundary layer transition location with a fast-response technique.

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.

This thesis presents an experimental study of the unsteady phenomenon known as transonic buffet. Completely developed in the High Speed Lab of the Aerospace Faculty at TU Delft (The Netherlands), the experiments were conducted in the transonic/supersonic wind tunnel TST-27. In order to achieve time and spatial resolution of the phenomenon, high-speed Schlieren and time-resolved particle image velocimetry were used as the main measurement techniques over two aerodynamic models, the NACA0012 and the OAT15A. The aim of the project was to achieve fully developed buffet conditions, to study the time evolution of the flow and the dynamic characteristics of the features involved in the phenomenon, to investigate the wave propagation in the flow that maintains the self-sustained character of buffet and to study the origin and dynamic development of these waves.

An spectral analysis of the shock wave motion was done over all the experimental test cases to analyze the influence of different aerodynamic parameters (Ma, Re, AoA) over the phenomenon and to come up with the cases were buffet is most fully developed. The results showed that, as expected, there is a range of values of these aerodynamic parameters under which buffet develops, and out of this range it just vanishes. Buffet was fully achieve for the OAT15A model at Ma=0.70, AoA=3.5º and fixed transition at x/c=7%, on the other hand, buffet about the NACA0012 was not fully developed, but it was found to be stronger at Ma=0.70, AoA=4º and free transition. From the PIV results, the instantaneous and the phase-average velocity fields showed how the actual flow development takes place and all the features involved in the buffet cycles. It was shown that under fully developed buffet conditions, massive separation takes place from the shock foot and extends beyond the trailing edge during the upstream motion of the shock wave, whereas reattachment occurs in that region during the downstream motion of the shock.

Finally, a correlation analysis of the velocity fields demonstrated the presence of the so called upstream traveling waves shed at around 2000 Hz and traveling upstream at -80 m/s. Furthermore, it was demonstrated that the vortices shed from the shock foot when full separation occurs are not the cause for the upstream traveling waves, instead they are originated by another type of disturbances shed when reattachment occurs at the shock foot. This made possible a reformulation of the hypothetical dynamic evolution of the features involved in the buffet cycles, explaining the links between them and the physical mechanisms that cause their self sustained development.

Laser diagnostic techniques such as Particle Image Velocimetry and Particle Tracking Velocimetry to measure the flow of a fluid around an object have been in prevalence for a few decades. Typically, the fluid, whose motion is of interest, is seeded with micron scale particles and illuminated with a laser. Due to scattering inefficiencies of these micron scale particles the maximum measurement volumes were about a few litres. Development of the neutrally buoyant sub-millimetre scale Helium-Filled Soap Bubbles allows for much larger scattering cross-sections thus enabling large-scale measurements in air (Caridi et.al. 2015).

Large scale Particle Tracking Velocimetry has been used to study the flow in the wake of a cyclist at typical time-trial speeds. A novel PTV algorithm known as Shake-The-Box (STB) (Schanz et.al. 2013) has been used to obtain the particle tracks from the images acquired in a thin volume in the cyclist’s wake using the wake-scanning method. The Lagrangian particle tracking technique was found to be more accurate than the Time Resolved Tomographic PIV through another experiment. Particularly for large field of views involving scanning at different positions, it produces whole-field results free from the boundary effects that arise between different positions of the scans. This yields accurate velocity components and their gradients that are crucial for pressure reconstruction.

Flow-field results from 4D-PTV show that the wake structure is similar to those found in the literature. Two large regions of momentum deficit are present, one behind the thighs and one near the lower legs and the wheel axis. Furthermore, they reveal the presence of some vortices which were not reported previously. Based on the observed signs of the vortices the points of origination of these vortices are identified based on separation mechanisms. There is a good agreement in the flow-field results with the literature. Pressure fields reconstructed from the velocity flow-field show low pressure pockets in the regions of the vortex cores and separated flow over the lower back.

Wake flow-field is utilised to compute the drag using the control volume approach (van Oudheusden 2007) and compared with the drag forces obtained from an external balance measurement to obtain the accuracy of large scale 4D-PTV. In total, five measurements at different velocities in the range of 13m/s to 15m/s are analysed. The accuracy of the drag estimates from 4D-PTV is obtained within 5%. Out of the three terms of the drag force, the momentum term contributes approximately 95% and the contribution from the pressure term is less than 0.3%. Rest is contributed by the Reynolds stresses in streamwise direction. Almost all the variation in the drag force comes from the momentum term.

A combined PIV and Schlieren measurements have been carried out in the transonic-supersonic wind tunnel (TST-27) to investigate the effect of exhaust plume and the variation in nozzle length on the flow topology and mean pressure distribution on the wake of axisymmetric backward facing step model at freestream Mach numbers of 0.76 and 2.20, respectively. Four different nozzle length configurations with and without the presence of a supersonic exhaust plume have been tested. Testing with different nozzle length configuration resulted in flow cases where the shear layer reattachment occurred on the nozzle (solid reattachment), on the flow downstream of the nozzle (fluidic reattachment), and intermittently on the nozzle and on the flow (hybrid reattachment). A qualitative identification of the effect of exhaust plume and the variation in nozzle length on the flow topology on the wake of axisymmetric backward facing step model, at the above mentioned conditions, is successfully described by means of Schlieren visualization. The topological flow features for the subsonic and supersonic flow cases has been identified (i.e shock waves, Prandtl-Meyer expansion fan, boundary layer, separated shear layer, recompression and plume shocks). On the other hand, a quantitative identification of the effect of exhaust plume and the variation in nozzle length on the flow topology and mean pressure distribution on the wake of axisymmetric backward facing step model, at the above mentioned conditions, is successfully done by means of planar PIV. The planar PIV measurements in the wake region of the model provided detailed information of the mean flow field properties (i.e mean velocity, turbulent kinetic energy, Reynolds stresses and reverse flow). Using these mean flow field properties data, the mean pressure distribution is reconstructed based on the momentum equation.It has been shown that an increase in nozzle length and the presence of an exhaust plume caused an increase in mean reattachment length at freestream Mach number of 2.20, while no significant change in mean reattachment length was noticed at freestream Mach number of 0.76. Significantly higher turbulent kinetic energy levels have been observed for L/D = 1.8 cases where solid reattachment occurred, at freestream Mach number of 2.20. In contrast to supersonic flow cases, the flow cases at freestream Mach number of 0.76 showed a significantly lower turbulent kinetic energy levels. Comparisons of flow cases with a long nozzle without a plume and flow cases with a short nozzle but with a plume suggest that the presence of the plume cannot accurately be modeled by replacing the plume with a solid geometry. From the pressure results it is observed that the location of the low-pressure region downstream of the base remained unchanged for different flow cases with and without exhaust plume and for different nozzle lengths. Furthermore, it has been shown that an increase in nozzle length leads to higher local pressure at the nozzle exit and hence results in a less under-expanded for the supersonic flow cases or more over-expanded plume for the subsonic cases.

Junction flow denotes the fluid phenomenon where the flow on a flat surface encounters a protuberance. This type of flow, which is highly three-dimensional, turbulent and unsteady, usually leads to the generation of a Horseshoe vortex at the obstacle's leading edge. Existing extensive experimental and numerical studies focus on the inception of such vortex, as it exhibits a bi-stable behaviour which persists downstream in the vortex legs. In a practical situation such as wing-fuselage juncture, the created vortex interacts with the flow around the aircraft. Consequently, the interference drag increases and the wake can reduce the effectiveness of the stabilizers. To tackle these effects, control devices are implemented to alleviate the strength of the vortex, being leading edge fairing the most widely used device.

In this research project the junction flow is subjected to the influence of different control devices: leading edge fairing, vortex generators and the novel antifairing are tested. The flow field is captured with the state-of-the-art large-scale tomographic PTV technique. The aim of this project is to characterize the vortical structures associated with each passive control device.

The time-averaged flow field results show the fairing is the most efficient device in reducing the presence of the Horseshoe vortex, but it is also the one which contaminates a larger wake area with the vortex wake. The vortex generators are the worst performer both in the mitigation of the vortex and the turbulence level in the wake. Nevertheless, the performance of such control devices highly depends on the location of the vortex generators. Thus, it is speculated that an optimal position exists which can yield better results. The antifairing shows minimal differences in the velocity field and the vortical structures compared with the reference case, but with a slightly lower turbulence level.

An auxiliary experiment with stereoscopic PIV is performed to the wake of the junction, in which the momentum deficit, a good indicator of the drag, is calculated for the different configurations. The leading edge fairing shows minimal drag reduction associated with the HSV of just about 1-2%. The VG once again have proven not being very effective with an increment of 4%. A reduction of more than 15% is measured for the antifairing case. Although the associated flow topology assimilates to the reference case.

This study investigates the HSV system around the junction as a volume. In comparison with most of the previous experimental studies on the subject, performed mostly with planar PIV and limited to just a few planes, the volumetric measurement can deliver more information with a single measurement. Therefore, it can better define the flow topology and vortical features and facilitate the understanding of the control devices working mechanisms.