Carbon capture and storage (CCS) has emerged as a cornerstone technology for mitigating climate change, as highlighted by the Intergovernmental Panel on Climate Change (IPCC) and International Energy Agency (IEA). A key component of CCS involves the safe and efficient transport of CO2 between capture sites and storage facilities, often over long distances. Pipelines can be an attractive mode of transport, with CO2 typically maintained in a liquid or supercritical state to optimize density and viscosity. However, rapid depressurization events, whether due to intentional releases or accidental pipeline ruptures, lead to complex two-phase flow dynamics. Accurately modelling these phenomena is critical to designing pipelines that are both safe and cost-effective.
This thesis investigates various one-dimensional, unsteady, compressible two-phase flow models to simulate CO2 depressurization scenarios in pipelines. All models are implemented using a finite volume method and discretization is done with the HLLC approximate Riemann solver, enabling the resolution of shocks and dis- continuities. Thermodynamic properties are computed using equations of state (EOS), with a focus on the Span-Wagner (SW) EOS, to capture the CO2 unique phase behaviour under high-pressure conditions. The models are validated against existing experimental data from SINTEF’s depressurization facility, which provided high-resolution measurements of pressure and temperature during rapid phase transitions.
Three different models are initially looked at: DF3, DF4, and TF5. Key findings demonstrate that while the DF3 model provides accurate predictions of pressure variations over time, it underestimates the initial pres- sure drop by up to 20 [bar]. In contrast, the DF4 model, through the manual adjustment of the mass transfer term (Γ) using the relaxation parameter ,θ , more accurately captures the initial transient behaviour, aligning closely with experimental data. Both models ultimately converge to approximately the same state after 20 [ms], with no more than 4 [bar] difference between the models.
Temperature predictions, however, pose a greater challenge. The DF3 model exhibits a spurious downward temperature spike immediately following pipeline rupture, while the DF4 model predicts an even greater initial temperature drop, neither of which align with experimental observations. The delayed cooling effect observed in experiments is not captured by either model, highlighting limitations in the energy equation and the need for additional source terms to account for temperature drop delays.
Limitations of the models include challenges in simulating temperature variations near the critical point and an inability to accurately model delayed temperature effects for both the DF3 and DF4 models. Recommendations for future work include developing a generalized mass transfer model, incorporating implicit numerical schemes for stability near critical conditions. Furthermore, the TF5 model shows promise for improving temperature predictions over longer timescales.

The 'hammerhead fairing' (HHF) launcher configuration uses a payload fairing that has a larger diameter than the launch vehicle. This allows for larger payloads to be carried using the same launcher components. However, the HHF leads to high aerodynamic forcing in the transonic regime called buffeting. If unaccounted for, buffeting can lead to launcher failure. The flow phenomena associated with buffeting were investigated on the Vega-E launcher, which is currently being developed by the ESA, using Pressure Sensitive Paint (PSP) in the TST-27 tunnel.

PSPs are a relatively new flow measurement technique that allows for quantitative measurements of surface pressure levels in a non-intrusive manner. High spatial resolution, which is equivalent to using millions of pressure taps, can be achieved using PSP. This allows for CFD-like surface pressure resolution to be obtained from wind tunnel experiments. However, PSPs had previously never been successfully implemented in Delft. Therefore, a test set up was devised for experiments in the TST-27. The final implementation was found to have a pressure uncertainty of at most 2.5 kPa, with a spatial resolution of 0.088mm.

First, validation measurements were done on the Coe and Nute Model 11. The Model 11 is a common HHF test case in literature that is used to study the buffeting phenomenon. Results obtained in the TST-27 were compared to historical data and showed good matching when accounting for blockage. After the validation, measurements were taken on the Vega-E launcher at different Mach numbers and angles of attack to quantify the effects of each variable on the pressure field. Lastly, the impact of protuberances on the model was studied. The protuberances that were used mimic those that will be present in the full-scale launcher.

The aerospace industry’s push for faster, lighter, and more durable designs highlights the critical challenge of panel flutter in supersonic flight, a poorly understood phenomenon intensified by shock-wave/boundary-layer interactions (SWBLI). This study experimentally investigates the effects of static and dynamic panel deformations on shock-induced panel flutter at Mach 2.0, using open, closed, and ventilated cavity conditions. Results show that cavity closure significantly influences panel response, with the open cavity causing greater oscillations due to strong pressure waves entering it, causing resonance behaviour; closing the cavity seems to filter out this resonance behaviour. Additionally, temperature changes and wind tunnel confinement were identified as factors affecting panel and flow behaviour. The results underscore the importance of understanding experimental conditions to accurately interpret fluid-structure interactions in supersonic conditions.

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.

A Tomographic Background-Oriented Schlieren (TBOS) technique is developed to aid in the visualization of compressible flows. An experimental setup was devised around a sub-scale rocket nozzle, in which four cameras were set up in a circular configuration with 30° angular spacing in azimuth. Measurements were taken of the overexpanded supersonic jet plume at various nozzle pressure ratios (NPR), corresponding to different flow regimes during the start-up and shut-down of rocket nozzles. Measurements were also performed for different camera parameters using different exposure times and f-stops in order to study the effect of measurement accuracy. Density gradients and subsequently two-dimensional line-of-sight integrated density fields for each of the camera projections are recovered from the index of refraction field by solving a Poisson equation. The results of this stage are then used to reconstruct two-dimensional slices of the (time-averaged) density field using a tomographic reconstruction algorithm employing the filtered back-projection and the simultaneous algebraic reconstruction technique. By stacking these two-dimensional slices, the (quasi-) three-dimensional density field is obtained. The accuracy of the implemented method with a relatively low number of sparse cameras is briefly assessed and basic flow features are extracted such as the shock spacing in the overexpanded jet plume.

TOP contoured nozzles, with large area-ratios, are commonly employed in rocket propulsion systems as they feature an excellent thrust-to-weight ratio. A significant shortcoming to this design is that, during the startup and shutdown transients of a LRE, the internal nozzle flow progresses through a series of overexpanded flow states - Free Shock Separation (FSS) and Restricted Shock Separation (RSS), which produce critical loads associated with SWBLI and asymmetric flow separation. Exacerbated by FSI, this operational phase is known to generate the highest vibroacoustic loads at which payload and vehicle structures are subjected to when the engine is operated at off-design conditions.

Motivated by the importance of understanding how the interaction between the developing flow and the vibrating nozzle walls has an effect on supersonic noise generation and propagation, this work has studied the effect that wall compliancy has on the vibroacoustic loading of TOP contoured nozzles. This is demonstrated by means of cold flow tests carried out in the High Speed Laboratories of the Delft University of Technology on a stiff-walled aluminum nozzle, which serves as a baseline test case, and on a urethane-based compliant walled nozzle. Tests are conducted under comparable flow conditions and test parameters are measured by means of acoustic and optical techniques. Simultaneous recordings are performed and include the nozzle-wall deformation, by means of stereoscopic tracking of tracers on the nozzle lip, the imprint of the near-field acoustic signature, by means of arrays of pressure-microphones, and Schlieren imaging of the jet plume.

Measurement data allows for a Fourier decomposition of the nozzle lip displacement and of the acoustic pressure field in azimuth. Reconstruction of the instantaneous plume development enables the identification of the main flow structures responsible for noise generation.

Comparison of results between the two test articles highlights a different spectral content and directivity pattern. Correlation between the structural displacements and the acoustic signal, together with the use of DMD, quantitatively aids the investigation of how FSI has an impact on the generation of an aeroelastic tone at 180 Hz. Findings suggest that its production is the result of the periodic thickening and thinning of the shear layer owing to the heightened flapping motion of the nozzle lip preceding RSS transition, driven by an intensified shock foot instability.

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.

The transonic regime is of great interest for designing aircraft wings because most conventional aircraft fly transonic. When flying in this regime, the wing of the aircraft tends to experience a phenomenon known as buffet, which is characterized by an oscillating shock wave. The surface pressure of a structure experiencing transonic flow is always changing; however, the average surface pressure can be computed. Pressure sensitive paint (PSP) is a flow measurement technique that determines the surface pressure and serves to compute its average. Using PSP, the average surface pressure of an ONERA OAT15A supercritical airfoil model can be determined in transonic conditions. This allows to study the effect of shock control bumps on the shock wave and surface pressure of the OAT15A model, as well as to determine the pressure coefficients of the airfoil.

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.

As the space industry continues to grow rapidly, the development of reusable launch vehicles has become crucial in achieving cost-effective and sustainable access to space. Payload capacity optimisation has been at the forefront of this effort, leading to a renewed interest in hammerhead or bulbous payload fairing (PLF) configurations. These designs feature a larger diameter in the PLF than in the rest of the launch vehicle, enabling the same structure to be used for large payloads. However, the transonic regime poses unique challenges for these PLFs due to their susceptibility to flow separation and strong pressure fluctuations.

This master thesis investigates the influence of nose and boat-tail geometries on flow phenomena, particularly shock wave generation, around hammerhead configurations in transonic conditions. The research was conducted through three distinct experimental campaigns in the transonic wind tunnel TST-27 at the high-speed laboratory of TU Delft, employing schlieren, oil flow, and Particle Image Velocimetry (PIV) techniques.

The study revealed notable insights into the aerodynamics of the hammerhead PLF configurations. Boat-tails set at five and 15-degree angles were observed to broaden the range of shock wave oscillations, introducing an additional shock wave that could occasionally merge with existing ones. Moreover, the conic nose design induced higher shock wave oscillations, while the bi-conic nose introduced an extra shock wave compared to the conic and ogive noses. The study also found that altering the nose shape while keeping the boat-tail constant, or viceversa, resulted in similar effects on flow dynamics.

These findings underscore the critical role that nose and boat-tail geometries play in shaping the aerodynamic behaviour of hammerhead PLF configurations. The results have implications for such configurations’ design and stability considerations in transonic conditions. It was observed that raising the Mach number heightened shock wave oscillations and flow detachment. The angle of attack disrupted model symmetry, primarily impacting reattachment patterns. Furthermore, the conic nose exhibited greater unsteadiness due to oscillations in the shock waves compared to the other geometries. The study also provided detailed insights into shock wave spectral characteristics and identified potential influences of pressure wave oscillations on shock wave behaviour. The conclusions lead to recommend designs with bi-conic nose, and avoid boat-tail angles around 15 degrees.

A series of numerical simulations was executed using the high-performance cluster at TU Delft's Aerospace Faculty. The primary objective was to gain deeper insights into the intricate dynamics of a sonic transverse jet in a supersonic crossflow with an oblique induced shock wave, especially concerning mixing. These simulations delved into the near and mid-field regions, illuminating how the position and intensity of an induced oblique shock wave affect the original jet in crossflow configuration.

In a simplified context, a box simulated the combustion chamber near the injector. A transverse jet was introduced through a lower wall hole, intersecting the incoming crossflow. Wedges on the upper wall generated induced shocks at varying distances from the jet exit. These experiments spanned different locations and strengths of shock impingements, all relative to the diameter of the jet's injection port hole.

The investigation aimed to comprehend how an additional shock wave influences mixing within the combustor. In the upstream scenario, the shock's proximity to the jet exit led to enhanced compression and the emergence of Kelvin-Helmholtz instability. Downstream impingement, on the other hand, disrupted the usual behavior of this instability. In both cases, the initial downward movement of fluid structures toward the wall was followed by ascent, expanding the region for crossflow-wall mixing downstream. Close-to-jet exit shock impacts intensified these effects, causing explosive disruptions and enhanced mixing.

Ultimately, the study revealed that closer and more intense shock impacts near the jet exit had a significant negative impact on volumetric energy, detrimental to combustion and propulsion potential. Optimal results were achieved when the shock impingement occurred farther downstream. The research also highlighted the need for improvements in simulation techniques and recommended further exploration of various parameters for comprehensive insights. This work serves as a foundation for future studies tailored to specific needs, promising more insightful outcomes.

This Master thesis consists in investigating the formation and behavior of vapor and ice plumes. These plumes can occur in icy moons of our solar system, such as Europa or Enceladus, which are widely believed to have a liquid water ocean beneath their crust. These plumes most likely consist of a sub- surface reservoir of liquid water fed by the ocean, a crevasse, and a vent at the ice surface. Under certain conditions, the liquid water starts evaporating and, due to the high reservoir pressure, the result- ing vapour starts flowing upwards through the crevasse. Due to the high pressure difference between the reservoir and the vent at near vacuum conditions, supersonic plumes are formed. This work stud- ies the reservoir conditions and plume physics, gathering different models available in literature and comparing them to available data, such as that from observations from Cassini spacecraft. First, a model capable of describing the condensation phenomenon is sought. The effects of the consequent release of latent heat to be absorbed back by the vapour flow are also studied. Grains will then start to nucleate and grow. These particles can either collide and stick to the walls or keep flowing mixed with the vapour. Some of the above mentioned latent heat can be absorbed back by the icy walls of the crevasse and generate more vapour by sublimation. Further, this project also aims to extend the model mentioned above so that it considers the effects of a fully multi-phase, multidimensional flow, checking for the effects of rarefaction and hence the limits of the continuum assumption. Finally, a similarity analysis is performed so that the influence of working with scales as different as the channels used in the laboratory at TU Delft or the real dimensions of the crevasses found in Enceladus can be fully tackled. Any progress on the ongoing investigations about the physical characteristics of these plumes could be crucial to deepen our knowledge on geological mechanisms in icy moons of our solar system. This, in turn, could trigger research on organic compounds present in these moons, perhaps even allowing for the existence of life. Further, the model used throughout this work can be applied to study power production devices where condensation might play an important role, such as steam turbines for light-water-cooled nuclear reactors or turbines proposed to be used in innovative organic Rankine cycle (ORC) configurations, natural gas supersonic separators, supercritical CO2 compressors for large-scale carbon capture and sequestration (CCS) and micro-nozzles.

The concept of a spaceplane, a launch system that is a combination of an airplane and a rocket, is one of the most promising concepts of sustainable access to space. Along its trajectory, it experiences most extreme conditions with frequently opposing design requirements, making it one of the most complicated challenges in aerospace engineering. This project aims at assisting in its design process by creating a KBE framework that combines a 3D geometry generation module, with automatic discretization and low-fidelity analysis using a panel method, ZONAIR. The results of the low-fidelity simulations are then correlated with CFD calculations using a multi-fidelity data fusion technique called co-kriging, to provide a good balance between high accuracy and low computational cost, making the framework applicable in the preliminary design phase. During the study of the accuracy of the proposed algorithm it was found that just 30 high-fidelity points is enough to generate surrogate models of aerodynamic coefficients with expected relative error below 5%.

During the start-up and shutdown sequence of rocket engines, immense forces are generated that are required to propel the payload into earth's orbit or outer space. The bulk of the generated forces are in the form of thrust. However, forces are also generated that act perpendicularly to the thrust direction. These side-loads are unwanted byproducts of the complex flow features present in the nozzle and as such, research has been conducted in the past to address this issue and discover their origin. Conventionally, research into side-loads has been conducted using subscale nozzles that don't deform as the flow passes through them. Due to the rigidity of these models, the nozzles don't deform to any degree that would cause it to interact with the flow. This fluid-structure interaction, present in full-scale rocket nozzles due to their sheer size, is therefore not taken into account when testing on subscale nozzles. To this end, a rocket nozzle testing facility was designed and built to investigate the fluid-structure interaction present in rocket nozzles using flexible subscale rocket nozzles. Due to the laborious nature of designing and manufacturing the testing facility, no working subscale flexible nozzles could be manufactured within the given time.

To continue building upon this previous work, subscale rocket nozzles were poured using various types of polyurethane. To cast these nozzles, multiple mould parts were manufactured. To guarantee a smooth surface finish for the inner nozzle wall, an inner mould was CNC machined. To enable the casting of multiple nozzles with different wall thicknesses, outer moulds were 3D-printed. After having determined the correct variant of polyurethane, Smooth-On's 'Smooth-Cast 66D', three different nozzles were cast on which tests were conducted. The nominal nozzle was designed with a wall thickness of 2mm, with two additional nozzles being cast with a wall thickness of 1.5mm and 2.5mm.

Having manufactured the nozzles, they were mounted to the test facility. The test facility consists of a stand connected to an air tank capable of containing air at pressures of up to 40 bar. Due to pressure losses throughout the valves, the maximum attainable nozzle pressure ratio, or NPR, amounted to slightly over 30. Various tests were performed during which the NPR was either altered during the test, done by manually opening or closing the control valve contained in the system. Further tests were conducted at a number of constant values of NPR at which excessive vibrations occurred to gain further insight into the flow phenomena present at these critical values of NPR. Lip tracking was done as well as schlieren photography. For the lip tracking, correction fluid was applied to the nozzle lip as well as spray painting the outer wall of the nozzle black. The area surrounding the nozzle was made dark such that the predominant feature photographed would be the circular nozzle lip. Once the images were taken, the nozzle lip was discretised in 180 points and was compared to an image of the nozzle taken when the wind tunnel was closed. The end result is a discretised function of 180 points displaying the lip deflection at each azimuthal location. With the help of the discrete Fourier transform, the dominant eigenmodes could be extracted from the nozzle vibrations in the form of their Fourier coefficients. In addition to lip tracking, schlieren images were taken simultaneously with the aim of visualising the flow downstream of the nozzle exit. The brightness of a set of pixels was tracked and their spectral make-up was determined.

Along the entire scale of NPR, two regions were found in which most of the vibrations occurred. The first region of activity, was purported to be linked to the transition of separation regime from free shock separation, FSS, to restricted shock separation, RSS. The second region of NPR, in which the deflection was often even more excessive than the first, is highly likely due to the presence of a recirculation bubble passing over the nozzle lip. Out of the first three eigenmodes, named the breathing, bending and ovalisation mode respectively, the ovalisation mode accounted for the largest proportion of the deflection by far. Based on these transient tests, further tests were conducted at a constant NPR of NPR = 21.7 and NPR = 27.7. These were chosen as all three nozzles showed large deflections at these NPR. For the nominal 2mm nozzle, two additional values were chosen of NPR = 22.7 and NPR = 27.5 to conduct further tests at. The energy present in each of the first three eigenmodes was determined by calculating the variance of the Fourier coefficients. The energy in the modes during the constant NPR runs were compared to the energy in the modes during the transient tests. These were obtained by determining the energy of an NPR-envelop spanning 1 unit of NPR centered around the NPR at which the constant run was performed. By and large, the energy storage in the first three modes remained similar when comparing transient tests with constant NPR runs. The bending mode showed the largest deviation from this trend, mainly when in RSS. Frequency analyses showed the nozzle to oscillate at rates predominantly within 130Hz-400Hz. Wavelet power spectra showed no clear trend between the NPR and the frequency with which the nozzles vibrated. Strouhal calculations further confirmed that NPR would not have a drastic effect on the unsteadiness frequency of the flow. The testing helped provide insight into the mechanical vibrations of nozzles of various wall thicknesses in different flow conditions. However, due to the absence of quantitative flow visualisation and measured flow unsteadiness frequency, no concrete link could be made to the aeroelastic coupling and the fluid's role in the forcing of the vibrations.

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 5δ_ν - 6δ_ν for the design Reynolds number of Re_τ= 1270. Direct force measurements are performed in the M-tunnel at a Reynolds number range of approximately 630 < Re_τ < 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 17δ_ν 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.

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 flows cause thin panels to flutter, which is characterized by high-amplitude self-sustained oscillations, increasing the risk of fatigue failure. Flutter is known to be exacerbated when a shock wave impinges on the panel, creating a shock wave/boundary-layer interaction (SWBLI) which promotes separation of flow and leads to increased aerodynamic and thermal loading on the panel. This novel fluid-structure interaction (FSI), known as shock-induced panel flutter, poses a risk to the structural integrity of components used in high-speed aerial vehicles, such as rocket nozzles and supersonic engine inlets.

Due to the complex coupled interaction between flow and structural dynamics in flutter, experiments provide a better option to investigate the underlying physical mechanisms, rather than computational studies, which usually have to compromise between computational cost and resultant accuracy. The ST-15 supersonic wind tunnel facility at TU Delft is employed to conduct experimental measurements of shock-induced panel flutter, using high-speed Schlieren imaging to capture flow structures and stereographic Digital Image Correlation (DIC) to record panel displacements in separate campaigns. Tests are done at Mach 2 with fully-clamped thin panels, and the flutter response is excited using different shock strengths and impingement locations.

In both campaigns, accelerometers are used to measure spurious vibrations around the wind tunnel test section. This helps reveal the existence of external vibrations inherent to the facility, which are also found to drive the frequency of the panel flutter: at 760-770 Hz without an impinging shock, and 605-640 Hz with an impinging shock. The latter frequency is detected in both - shock motion and panel displacements – which establishes coupling between flow and structure despite measurements being non-simultaneous. Changing the shock impingement location does not have a significant effect on the degree of flow separation caused over the thin fluttering panel, which is always higher than the separation on a rigid plate at the same shock strength, thus proving that fluttering panels are not viable means of shock-induced separation control. A stronger impinging shock produces increased shock-induced flow separation but results in less energetic panel flutter, which is attributed to the higher post-shock pressure rise suppressing the panel. Flutter is found to be most energetic, and consequently, the panel is most susceptible to fatigue failure, when the shock impinges at 60% of the panel length.

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.

Supersonic panel flutter is a self-excited vibration, where energy is transferred into thin flexible panels, such as skin panels in supersonic vehicles, from the (supersonic) flow around them, which can lead to their rapid fatigue failure. Relevant examples of this include the failure of skin panels in the X-15 vehicle, or violent oscillations in non-adaptable nozzles during start-up, such as the space shuttle main booster. With the contemporary trends of decreasing weight and increasing operational speeds in the design of next generation aircraft and launch vehicles, such structures will become more common, and flutter needs to be inherently accounted for in the design.

In recent years, advancements in numerical techniques, such as Computational Fluid Dynamics (CFD) and the Finite Element Method (FEM), have led to numerous studies that address the effect of non-linearities, such as flutter behaviour in the transonic and hypersonic regime. In spite of that, experimental data needed to validate these studies is limited, based on coarse pointwise measurement techniques, and only focused on the structural response of the panel rather than on the entire Fluid Structure Interaction (FSI). The use of simultaneous, full-field and non-intrusive measurements can be used to address these issues. To this extent, panel designs were developed and an experimental campaign was proposed where Digital Image Correlation (DIC) and Schlieren are simultaneously used by means of high speed cameras to study classical panel flutter at Mach 2 inside the ST-15 wind tunnel at the Delft University of Technology. Additionally, a Laser Doppler Vibrometer (LDV) was used to validate the DIC measurements. The experiment allowed for a comparison with well understood theory, such that the setup can be proven to be robust, and can therefore later be used to assess the effect of panel flutter non-linearities or other supersonic FSIs.

Synchronisation between DIC and schlieren was demonstrated through a matching frequency spectrum, and maximum correlation inside corresponding pixel windows between panel displacement and schlieren grey-scale fluctuations with a zero time-lag. One finding was that during the panel down-stroke phase, a travelling wave character was observed which has previously been found only under transonic conditions. Furthermore, at moderate dynamic flutter pressures, a typical second bending panel flutter mode shape was observed, which grew into a single bending mode again with increasing higher dynamic flutter pressures. From the spectral analysis it was found that all configurations fluttered at the same frequency of 770 [Hz], independent of dynamic flutter pressure or edge conditions. This was higher than expected from analysis. These combined observations led to the conclusion that, in the current facility, a frequency lock-in occurs due to merging of the flutter phenomenon with a wind tunnel resonance vibration.

These observations provide an insight into combining these experimental techniques to study supersonic FSI behaviour, and particularly into certain unexpected behaviour encountered at the TU Delft ST-15 wind tunnel. This can be used to further tune future experiments and guide researchers to obtain a better understanding in the non-linearities found in panel flutter.

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.