Contrary to conventional wisdom favouring smooth surfaces for drag reduction, the last decades have brought forth different textured surfaces showing drag reduction in turbulent boundary layers, of which riblets are the most studied. Sirovich and Karlson (1997) introduced a different textured surface named chevron-shaped protrusions, indicating a drag reduction of 10% in turbulent channel flows. Later studies questioned their efficacy as no credible reproduction of the results was obtained throughout the years. The last credible reproduction study by Carrasco Grau et al. (2023) suggested that the reasons for this discrepancy could reside in the difference in model and test section size. All reproduction studies were performed in facilities with external boundary layers instead of internal and with a covered area no longer than 0.8 meters, instead of the 8-meter-long channel flow which was fully covered in these chevrons from the original study. This work will investigate just that, using a channel flow facility with dimensions more akin to that used by Sirovich and a model size of 2.4 meters. The study assembles and characterises an improved channel flow facility with dimensions identical to those used by tay et al. (2011) at the National University of Singapore (NUS), measuring a total length of eight meters and a test section of 2.4 meters. This facility used an array of 29 static pressure taps to determine the skin friction via the mean pressure gradient method, and utilised hot wire measurements in the midpoint of the test section for investigation of the flow mechanics of the chevrons. The measurements performed on a flat plate were compared to the results obtained at the NUS to characterise the facility and validate its suitability for single-point drag measurements in the order of 5-10% increase or decrease. Once this was confirmed the different configurations of chevron-shaped protrusions were investigated. The key findings of this study are twofold. Firstly, the developed channel flow facility at DUT proves proficient in generating canonical boundary layer profiles, enabling accurate skin friction measurements of textured surfaces. The facility provides good hot-wire measurements without sensor vibrations capable of investigating the boundary layer characteristics in this facility. Secondly, the study establishes that chevron-shaped protrusions are likely unsuited for reducing turbulent skin friction in turbulent channel flows. Despite an inability to definitively disprove the working hypothesis, the observed increase in drag suggests that the technique's efficacy is not determined by the facility type and model size. It is not considered worthwhile to conduct additional research into chevron-shaped protrusions as a potential technique for reducing drag in turbulent boundary layers. However, the minimal drag penalty associated with this technique opens new possibilities for the application of this technique in the aviation sector, mainly in aiding with separation control and heat transfer.

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.

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.

In the last decades, the prevailing belief that smooth surfaces offer the lowest drag has been challenged often. Scholars have, for example, introduced rough and modified surfaces to reduce turbulent skin friction. One of the technologies proposed in the literature is an array of chevron-shaped protrusions; however, there is no academic consensus on the drag performance of this technique. Furthermore, although the theoretical working mechanism has been documented well, there is no experimental evidence in the literature to support the hypothesis around this mechanism. In this study, the experiments from the literature are replicated, and new array configurations are tested to characterise the effect of individual parameters on the drag performance. The test plates are manufactured by applying vinyl protrusions of roughly 100 μm in thickness to an aluminium base plate. This thickness corresponds to 5δν - 6δν for the design Reynolds number of Reτ= 1270. Direct force measurements are performed in the M-tunnel at a Reynolds number range of approximately 630 < Reτ < 1850 to determine the drag performance. Furthermore, the coherent structures are characterised by means of 2D-2C PIV of a wall-parallel plane at a minimum distance of 17δν from the wall. The drag reduction reported in the literature could not be replicated, and the balance measurement results offer relevant insights into the drag performance of chevron-shaped protrusions. The results consistently show that the added roughness due to the presence of the protrusions is not the only parameter that determines the drag performance, confirming a meaningful interaction between the protrusions and the flow. Moreover, the results are found to be highly sensitive to the randomisation of the array. No substantial effect of the protrusions on the coherent structures close to the wall has been observed. In particular, no evidence has been observed that supports the working mechanism as proposed in the literature. An analysis of possible causes to explain the discordance between this study and the literature is performed. Based on this analysis and the results from the aforementioned parametric study, an improved design is proposed, and recommendations for future research are postulated. This technology has inherent benefits for real-world implementations as it can easily be (retro)fitted to aircraft by means of a foil. Further research into this flow control technique is thereby deemed relevant due to the combination of the large drag reduction reported in the literature, the advantages in practical applications, and the novel opportunities for additional investigations.

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 Reτ, 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 Reτ 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.

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.

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.

Current-day space vehicle are equipped with a passive Thermal Protection System (TPS), but these systems operate at the edge of their capabilities. The research into the next generation reusable space vehicle demands the use of active TPS, these systems are not limited to the capabilities of the used material but depend on the active cooling mechanism. A system of cooled metallic TPS is developed at the TU Delft resulting in an increase in the maximum allowable thermal load for a given TPS temperature, this system is the enhanced radiation cooling TPS. The research question is: "Can an active thermal protection system improve the flying characteristics and the flight performance of a re-entry vehicle?". The first step in answering this question is to verify the process behind the active TPS. This thesis aims to design an experiment to verify this process. To answer the research question two tools were developed: a flight simulator and a thermodynamic analyser. The flight simulator provides insight in the behaviour of the rocket and shows the consequences of implementing the enhanced radiation cooling TPS on the rocket. The thermodynamic analyser is used to estimate the wall temperature of the system and shows the effect of the sensitivity analysis of the design of the cooling subsystem. The enhanced radiation cooling system is designed to be mounted on the T-Minus Dart and an experiment to verify the impact of the active TPS is derived using the nominal mission of the rocket. Sensitivity analysis resulted in thin outer shell and a insulation sphere filled with the coolant. The outer shell needs to be as thin as possible in order to maximize the cooling effect. For an outer shell of 0.5 millimetres, the maximum allowable thermal load can be doubled in comparison with an non-cooled flight. If an identical thermal loading is assumed for two return vehicles, the active cooled vehicle can be designed with a nose radius of around 30 % smaller that the non-cooled vehicle. This will result in a vehicle with a higher lift-over-drag ratio, extending the return trajectory which increases the safety of the re-entry flight. The validity of performing an enhanced radiation cooling experiment is shown by the use of the tools developed in this thesis. The incoming heat flux can be doubled while maintaining the same wall temperatures if the cooling system is applied. This opens the possibility of redesigning re-entry vehicles to achieve a high lift-to-drag ratio.

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.

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.

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.

Curved exhaust ducts are used in aero engine applications for different purposes, including thrust vectoring, shielding of parts from the exhaust or improving stealth properties. Their integration, however, regularly represents a design problem due to flow separation and high aerodynamic losses occurring in the bend. Curved duct flows for both compressible and incompressible conditions have been studied extensively in the past. However, no experimental results of a high Reynolds number flow through a turbine exit guide vane (EGV) followed by a curved duct have yet been published. An in-depth CFD analysis of the aerodynamic effects is therefore carried out, using the RANS solvers TRACE (at MTU Aero Engines) and SU2, to analyze the flow field and geometric sensitivities of a high Reynolds number flow through an EGV followed by a 90 degree bent duct. The geometry of interest is investigated at a Reynolds number of Re=10^6 and a ratio of bend radius to duct diameter of one. The inflow conditions are prescribed to closely resemble typical flow conditions at the low pressure turbine exit plane of a turbojet engine. After validating the solver with experimental data using the test case of a 90 degree bent duct, an initial CFD analysis of the combined exit guide vane geometry with curved duct is carried out to identify the dominant flow phenomena and mutual effects of the EGV and the bend. Three zones of flow separations are found, each at the convex and concave sides of the bend and one at the lower side of the plug. Secondary flows caused by the bend are found to have an effect on the flow upstream of the EGV, leading to a non-uniform flow inlet angle and aerodynamic loading of the individual blades. Separation downstream of the EGV is influenced by the presence of low velocity wakes from the EGV. Subsequently, a sensitivity study is carried out to find the effect of different geometrical parameters on the flow field. Main investigated parameters are the ratio of bend radius to diameter (R_c/D) and the distance between EGV and bend (l/D). Additionally, the aerodynamic effects of the circumferential EGV positioning, swirl and the plug shape are investigated. It is found that an increase in both bend radius and distance between bend and EGV improves aerodynamic efficiency, while swirl can decrease pressure losses in the duct for small bend radii of R_c/D<0.8. Improving the plug shape and rotating the EGV allows to further increase the aerodynamic efficiency without weight increase. For further optimization of the geometry, it is recommended to include duct geometries with a non-constant bend radius and outer duct diameter to increase the design space.

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 0o, 15o and 30o, 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.

Drag in pipelines is composed almost solely of skin friction drag. The most common technique to achieve drag reduction (DR) is by adding drag-reducing agents. However, in the aerospace industry, various impressive passive drag-reducing techniques have been suggested to reduce skin friction drag in the past decades. Among these techniques, dimpled surfaces form a relatively unexplored terrain. Research into reducing skin friction drag in the turbulent regime by using dimples has been performed in the aviation industry since the '80s. The excessive amount of skin friction drag in pipelines forms an intriguing challenge to break new grounds. Several researchers investigated the potential of DR of turbulent flows using dimpled channels. Even though most of these studies that found DR are disputed, studies published at the National University of Singapore (NUS) obtained positive results time and again. NUS's exact test setup was reconstructed at Delft University of Technology, which has not been done yet as far as the author of this report is aware. Identical pressure measurements were performed, yielding an absolute DR of ≈ 5%, which is slightly less than what was obtained at NUS (>7%). Flat plates return a DR between 8-15%, dimpled test plates returned at DR between 12-20%. The test plates were covered for 99.5% with diamond-shaped dimples of 100 mm long and 50 mm wide. In total, 29 pressure taps were used to determine the change in drag in an 8 m long channel of 20 mm in height. Tests were done at Reynolds numbers, based on half channel width and centerline velocity, between 6,000 and 40,000. Accurate results were perceived up to a Reynolds number of 21,000, likely due to test-setup limitations. Multiple verifications such as two-dimensionality of the ow, comparison of theoretical and experimental skin friction coefficients, and instantaneous pressure tests were used to allow for an objective analysis. The pressure measurements were also supported by 1D hotwire anemometry (HWA) tests and surface oil ow visualizations (SOFV). The majority of the investigated boundary layers were absent of anomalies. It should be mentioned that the viscous sublayer could not be captured, neither a quantitative momentum analysis was performed. However, a shift in the velocity profile, acquired at the same test location, for different test plates was observed. Furthermore, an increased velocity near the wall was observed for dimpled test plates. Surface oil ow measurements did show similar ow patterns to what was recorded at NUS. However, the actual behavior is not investigated through HWA. Hence, it cannot be confirmed nor denied if the change in drag is a consequence of these near-wall ow mechanisms. SOFV did not show irregular ow structures such as ow reversal. Finally, tests were performed with a correction for test volume increase caused by the dimples. After this correction, the drag over dimpled plates increased instead of reduced. Considering other studies on dimpled surfaces that also found an increase in drag, it is strongly believed that the positive effect of dimples in turbulent channel flows does not stem from skin friction drag but rather from the increase in channel volume. Although the precision of the results was relatively high, the accuracy of the wind tunnel was insufficient. Additional research is required to narrow down the 8-15% error that was obtained while testing at plates.

A loss of vacuum accident in the hyperloop tunnel will likely occur during its operation, and it may have a severe impact on its operational safety. Nevertheless, the hyperloop tube breach has only been researched superficially. Two aerodynamic phenomena due to the breach are identified: a (quasi-)steady underexpanded jet and an unsteady blast wave. The overpressure due to the confined blast wave impacting the hyperloop vehicles is determined using various methods. Numerical simulations of the breach applying the 3-dimensional Euler equations show that quasi-1-dimensional (Q1D) approaches only suffice for smaller holes. Although their analytical implementation is less involved than numerical analyses, the Q1D models underpredict the blast wave overpressure, because they cannot account for energy input after the critical shock formation. The (pressure) drag on a vehicle inside the evacuated tube may increase more than an order of magnitude due to the blast wave emanated from the breach location.

Supersonic intakes are essential in assuring proper combustion for operating ramjet engines on cruise missiles. However, due to increasing threats and a rapidly changing service environment, today's intakes do not fit anymore to comply with the new set of requirements to tackle the above challenges. Therefore, a novel supersonic intake design is investigated to evaluate the possible integration of this design on modern cruise missiles. The Submerged Supersonic Intake is a novel intake design that flushes with the external missile hull. An experimental model was investigated in a blow-down wind tunnel with a Mach 2.0 freestream flow. The analysis revolved around the occurrence of the pseudo-shock wave inside the intake's constant-area duct. The flow domain was investigated using Particle Image Velocimetry, Schlieren imagery and wall pressure measurements. Submerged intakes show comparable behavior to generic intakes and the Total Pressure Recovery resembles the results found from experiments in open literature.

Pressure reconstruction based on particle image velocimetry (PIV) refers to the determination of pressure data from pictures of small tracer particles added to a flow. The technique possesses a unique combination of beneficial characteristics in that it non-intrusively provides simultaneous pressure and velocity data in the flow field without the need for instrumentation or other preparation the model. The present research provides a structured overview of different approaches to PIV-based pressure reconstruction and characterises their (relative) performance, particularly when applied to a transonic base flow. The unsteady, large-scale behaviour of this flow constitutes is subject of active research in the context of launcher aerodynamics to which experimental pressure field data would make a valuable contribution. Two techniques are analysed in depth through theoretical analyses, a simulated experiment based on a numerical simulation and several wind tunnel experiments: pseudo-tracking for the determination of instantaneous pressure fields and the Reynolds-averaging approach for the determination of time-averaged pressure fields. Various PIV-based methods for instantaneous pressure determination are capable of reconstructing the main features of instantaneous pressure fields, including methods that reconstruct pressure fields from a single velocity snapshot. Highly accurate pressure fields can be obtained by tracking individual particles in combination with advanced processing techniques. In view of this outcome, it is recommended to let the choice for a specific technique be guided by the desired accuracy, resolution and dimensionality of the pressure results, while taking taking into account practical considerations, in particular limitations in the capabilities of available measurement equipment and the complexity of the measurement system. Without such intent, the potential difficulties and complexity of data acquisition were demonstrated with the use of a 12-camera/2-laser PIV system. For instantaneous pressure reconstruction through pseudo-tracking new insights were obtained on its spatio-temporal filtering behaviour and the propagation of velocity measurement errors. A cut-off peak-response is specified as a function of the temporal track length and spatial resolution. Novel approaches are suggested to determine suitable temporal track lengths on the basis of the variation in material acceleration with track length and on the basis of pressure power spectra. Such spectra were also used estimate the local error margin of reconstructed pressure values. For the implementation of pseudo-tracking, it is recommended to first construct tracks by a combination of a second-order integration method and linear interpolation, using an integration time step that is sufficiently small to meet the Courant–Friedrichs–Lewy condition. The material acceleration may subsequently be estimated from the tracks by means of least-square fitting of a first-order polynomial or central differencing depending on the type of input data. When calculating mean pressure fields with the Reynolds-averaging approach, it is recommended to only include the terms that are associated with the mean flow and Reynolds stresses. The impact of neglecting spatial and temporal density variations may be estimated as the difference between pressure solutions calculated with and without density-gradient terms. After validation, the approach was employed to study the effects of an exhaust plume and nozzle length on transonic and supersonic axisymmetric base flows. Amongst others, the results showed that depending on the nozzle length the presence of a plume may cause a decrease in base pressure in the transonic flow cases and an increase in base pressure in the supersonic flow cases, indicating the effects of entrainment and displacement, respectively. The results furthermore highlight the need of considering during vehicle design, that a longer nozzle in which a plume expands further, not only corresponds to a lower exit pressure in the plume, but also to a different ambient pressure near the nozzle exit.@en