Desert sand acoustic emissions are produced when a “sonic sand” is sheared locally or by a natural dune slipface avalanche, resulting in a brassy sound between 50 and 400 Hz. This type of sediment exhibits particular granulometric, shape and surface characteristics, due to the grains’ erosion and transport history, and emits sounds when the sheared grain layer vibrates in a synchronized manner, much like the membrane of a speaker. Recording such sand acoustic emissions on Mars (and perhaps other planetary environments) using rover microphones could thus become a new form of observable for scientists to estimate the surface sediment’s characteristics and history from a distance, but also the granular flow dynamics taking place. To determine whether this approach could be viable in the future, it is essential to evaluate how the Martian environment may affect sand acoustic emissions differently than on Earth. After showing that the muted Martian soundscape would likely allow rovers to detect such signals from a few tens of meters, the present thesis studies the impact of the interstitial air pressure within the sand bed on the sound emission mechanism of such sonic desert sands.

In this project, silent and sonic desert sand shear flows are induced under a range of pressure levels, from terrestrial ambient pressure to Mars-like pressure, within two separate, manually operated vacuum chamber setups: a smaller chamber shaken to create the sounds, and another longer chamber that better replicates avalanche-like sand flows. The motion applied and sound produced are measured using an accelerometer and a microphone inside the chamber. Metrics in the time and frequency domains are defined to analyse the changes in sound energy, amplitude, and frequency components produced at different pressure levels. Firstly, the silent sand tests are used to establish how the air pressure level within the experimental setup affects the regular sound of sheared sand (i.e. grains impacting one another) and more generally the sound emission of “normal” sounds, whose emission mechanisms do not depend on grain packing and synchronized motion. Then, a simplified theoretical model of how the sound pressure level (SPL) of a sound evolves with decreasing acoustic impedance, is derived and validated using the silent sand measurements performed. Finally, the sonic sand measurements are compared to the SPL model and silent sand measurement results, which are used as a baseline for nominal sound production behavior, to evaluate how the interstitial air pressure affects the amplitude and signal energy of the sheared sonic sand emissions. Furthermore, differences in the sand acoustic emissions’ frequency spectra and time duration across pressure levels provide information about the possible physical changes occurring in the granular flow dynamics of the sheared sonic sand.

In both experiments, the dominant frequency very closely follows the trend of the motion metrics used, as described in the literature, and remains very consistent across pressure levels. This suggests that the maximum sheared sonic sand layer thickness is independent of the interstitial air pressure. Then, the sonic sand emissions see an increase in the sound amplitude and signal energy related metrics from ambient pressure to 413.25 mbar, unlike the gradual decrease predicted by the SPL model and the trend of silent sand measurements with decreasing pressure. Below 413.25 mbar, the results suggest a stabilized behavior, with the acoustic metrics of the emissions following the model. Furthermore, in the avalanche-like emissions, a new frequency component slightly higher than the dominant frequency emerges as the chamber pressure decreases. These observations are evidenced in the time-domain, where the sand acoustic emissions seem to initiate earlier in the granular flow at 413.25 mbar and below, resulting in greater acoustic pressure levels being produced, compared to those at terrestrial pressure. It is hypothesized that more sheared sonic sand grains synchronize at 413.25 mbar and below (compared to terrestrial air pressure), and thus increase the amplitude of the sound wave produced. For avalanche-like flows, the new frequency component that appears with decreasing pressure level seems to suggest that the minimum sheared layer thickness threshold required to produce an emission is lowered at lower pressure, which leads to a higher frequency produced initially until the full layer forms, ultimately decreasing the frequency. Further research is required to confirm these preliminary findings and theories.

Interest in powered flight on Mars has been growing in recent years, and as such new developments are being made in determining the effects the Martian atmosphere has on the flow physics and performance, specifically with regards to similarity parameters: the Reynolds number & the Mach number. This paper aimed to investigate whether the onset of transonic flight could result in decreased separation while improving aerodynamic performance due to the presence of a smeared shock. 2-D URANS simulations of a NACA-0012-34 airfoil were conducted with a SST k-w, y-intermittency model. The results find that an expansion fan delays separation, while the smeared shock weakens the adverse pressure gradient, similar to a shock control bump. The result is a minute increase in the lift-to-drag performance. The development of coherent structures within the separated shear-layer have also been seen to be impacted due to the stabilizing effect of compressibility on the Kelvin-Helmholtz instability.

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.

This work presents an experimental and analytical investigation of unsteady loading of a low Reynolds number propeller and its relation to aeroacoustics. The propeller is positioned at an angle of attack with respect to the freestream which causes a variation in loading along the azimuth. The propeller is operated at 4000 RPM with an advance ratio of J = 0.4, corresponding to a free stream velocity of 8 m/s. The chord-based Reynolds number of the propeller is of the order of 10⁴ and is investigated at two angles of attack: α_r = 7.5° and α_r = 15°. For reference, a steady case, α_r = 0◦, is presented as well. Loads and noise measurements were carried out in the A-Tunnel of TU-Delft. Two microphone array configurations were used to mimic the observers positions above and underneath the flightpath of the propeller. Underneath the flightpath, an increase in SPL was found with increase of the rotor angle of attack. Above the flightpath, the opposite effect was observed, resulting in an asymmetrical directivity profile when the rotor is at an angle of attack. A quasi-steady loading model, based on the lifting line code Xrotor, is used to determine the azimuthal variation in loading. The variation in tip speed along the azimuth due to the propeller angle of attack is prescribed to determine the variation in loading in one rotation. Far-field loading and thickness noise are computed using an aeroacoustic model based on the Ffowcs-Williams & Hawkings’ equation to predict the harmonic noise in the frequency domain of a rotating dipole and monopole, respectively. The numerical predictions for the steady case show good agreement for both the aerodynamic and aeroacoustic performance when compared with loading and noise measurements. For the α_r≠ 0° cases it was shown that unsteady loading noise increases the SPL for all harmonics, with a larger contribution at higher harmonics.

The increasing importance of climate change and stricter regulations for greenhouse gas emissions raise the pressure on the automotive industry to develop more efficient products. Besides a reduction of emissions by new combustion engines, electric vehicles are being introduced to lower the fleet fuel economy. Especially this new drivetrain technology benefits from an improved driving resistance of the vehicle, as an increase in range is highly valued for current electric vehicles. The aerodynamic drag, a considerable contribution to the driving resistance, therefore plays an important role in automotive development. With a significant contribution to aerodynamic drag, the wheels and surrounding wheelhouses are considered to have potential for possible improvements and are currently one of the focuses of research. This thesis examines the aerodynamic flow in a rear wheel arch of a low drag electric sedan to identify areas of potential improvement and derive recommendations for future developments. Three major parameters of the rear wheelhouse geometry are therefore modified and their influence on the aerodynamic drag and the flow field is assessed. These include the wheelhouse radius, wheelhouse width and the wheel track. For this purpose, Reynolds Averaged Navier Stokes (RANS) simulations are used to both calculate the effect of the variations on aerodynamic drag and analyse the origin of changes in the flow field. For parameters where it is possible, tests on a detailed full scale model are carried out in a modern full scale wind tunnel.

Airfoils in deep stall have been a subject of extensive computational discussion in the past, with multiple efforts being performed by various institutions to test their solvers and turbulence/hybrid sub-grid scale (SGS) models for their use in massively separated flows. This case has an important application in turbomachinery where unsteady flow is encountered - including wind turbines and helicopter rotors - for studies in both aerodynamics as well as aeroacoustics. Stalling conditions are caused in these applications when there is blade-wake interaction, and also in the presence of strong gusts. The change in the local flow physics in both these cases is caused by a local variation in the angle of attack, which can affect the aerodynamic performance from entire blades.

Although there have been studies in the past that have involved various hybrid SGS models and turbulence models to investigate the effect of them for this flow case, the shear-layer adapted SGS (SLA-SGS) model has not found precedence for the purpose of aerodynamic investigation in this flow case. Therefore, this thesis provides an extensive aerodynamic investigation of an airfoil in deep stall, with a comparison of flow field visualization results using this approach, and comparing aerodynamic performance parameters with other hybrid SGS models. An attempt has also been made to compare the simulation results with other turbulence models and meshing strategies for similar airfoils, to present a case to estimate the similarities and differences in performance.

The results show that the there is an overprediction of the trends obtained from the 3D results compared to the 2D experimental data. The sectional 2D results also show an overpredicted output compared to experimental data. Contradictory results are obtained in comparison to the reference literature used for different SGS models.

Finally, additional studies have been performed on the aerodynamics points of view, with an overview of the acoustics code implementation in SU2 being provided. This is in order to provide an insight into not only the working of the code, but also the possibilities of implementing the code for providing acoustic outputs for the present case.

In the global search to more sustainable developments, research in the field of propeller propulsion systems is experiencing a renewed interest. The combination of efficiency and noise production is not studied as extensively as propeller efficiency alone in past studies. In this study, the relation between propeller noise and efficiency is investigated. Adjusted parameters are sweep, advance ratio and the collective pitch angle. The performance of randomly generated blades is evaluated using both RANS and BEM methods first, after which the noise is evaluated using Hanson’s frequency formulation. All results are combined in a multi-fidelity kriging surrogate model. The results suggest that maximum efficiency and minimum noise cannot be obtained at the same operating conditions. Adjusting advance ratio and pitch of a given propeller during flight may reduce the noise emissions at the cost of efficiency. Sweep is also shown to have a considerable effect on both efficiency and noise.

Wingtip-mounted propellers have the ability to enhance the propeller-wing system performance. For inboard-up rotating tractor propellers, the swirl in the propeller slipstream can be used on the wing to reduce the influence of the wingtip vortex, decreasing drag. Furthermore, lift is enhanced by increasing dynamic pressure and angle of attack locally. In this thesis the performance of a propeller-wing system with wingtip-mounted tractor propellers is investigated using a low-order numerical model. The numerical model is used to investigate the influence of the main design parameters for wingtip-mounted propellers. Response surface analysis is used to predict the propeller performance throughout the design space, while minimizing the number of evaluations needed. This has led to insights in the influence of wing and propeller design on the performance of a propeller-wing system.

With the goal of decreasing the environmental impact of aircraft, ducted propellers emerge as an efficient propulsive alternative. Ducts are able to increase the thrust to power ratio of a propeller system by both producing thrust and/or lowering tip losses of propellers. In this thesis, RANS CFD simulations were used to analyse the possible impact of modifying a propeller duct shape from a circular to a square geometry. Initially, the two duct designs and the propeller were studied as isolated cases, in order to characterise their aerodynamic performance. In the installed simulations, the propeller was first modelled as an actuator disk, and afterwards with a full blade model. The results indicate two main disadvantages of square ducts. Square duct corners were found to be prone to separation, and to contribute towards the generation of strong vortices. This research can be important for the future study of unconventional ducted propellers, for example with applications to distributed propulsion concepts.

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.

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

One of the major challenges that engineers face when designing a high speed vehicle is aerodynamics heating. While ideally the external surfaces are expected to maintain a smooth profile, this is not the case in most vehicles, which might be laden with discontinuities like steps, gaps, and other protrusions. These could alter the local flow field and could have a drastic impact on the heating and hence should be properly studied. In this thesis, the effect of large protuberances, when placed in both a laminar and turbulent boundary layer is experimentally investigated. Cylindrical elements with height greater than or equal to the boundary layer height, placed on a flat plate were used as protuberances. Quantitative Infrared Thermography was used as the prime investigative tool, assisted by high speed Schlieren and oil flow visualisations. For the laminar interactions, a series of high and low heat flux regions were observed in the wake of the protuberance due to the presence of stream wise vortices. A local peak due to flow reattachment was found downstream of the cylinder along the centreline, whose length from the cylinder trailing edge scaled with the height of the cylinder and was nearly invariant with the diameter and the flow unit Reynolds number. The span wise distribution of Stanton number revealed that the location of heat flux peak formed due to the symmetry plane counter rotating vortices, scaled with the diameter of the cylinder with their peak heating increasing with increasing in diameter. This revealed the possible dependency of the strength and location of the symmetry plane vortex pair with the diameter of the protuberance. The flow transition induced by the elements considered here, moved upstream with increase in unit Reynolds number. Moving upstream, a presence of span wise vortex system was verified. A mean separation length was measured from the centreline heat transfer distribution and an empirical formulation was presented and shown to have good agreement with the current dataset along with recent literature. For the turbulent interaction, direct heat transfer measurements were acquired downstream of the protuberance for the first time. A centreline high heat transfer was observed, with heating of $1.5 - 2$ times the local turbulent case. A series of high and low heat flux regions were observed in the wake, dictating the possible presence of stream wise vortices. Similar to the laminar case, the location of reattachment heating was shown to scale with the height of the cylinder. Additionally, span wise heat transfer profiles show that the wake width increases with increase in height, while remaining fairly constant for a given condition, throughout the length of the measurement domain. Time resolved Schlieren images captured upstream of the protuberance showed a very unsteady separation shock and a relatively steady bow shock wake, together forming the lambda shock system. Probability density functions generated for the upstream separation length revealed that both the mean and the standard deviation of the separation length increases with increase in height of the cylinder.