This research investigates the application of acoustic excitation as a novel method for transverse forcing to achieve friction drag reduction in a turbulent boundary layer. Traditional transverse wall motion, while effective in reducing drag, poses experimental challenges due to mechanical complexity and limitations at high Reynolds numbers. This study explores an alternative approach where transverse velocity gradients are induced through oscillatory acoustic fields rather than wall motion. An experimental setup was developed at the Delft University Boundary Layer Facility (DUBLF), incorporating a system of phase-synchronized speakers to generate controlled transverse acoustic forcing. Particle Image Velocimetry (PIV) was used to characterize the boundary layer and assess the effect of forcing across a range of Reynolds numbers. Results show measurable reductions in friction drag, with the optimal configuration achieving up to 6.02% drag reduction at Reτ = 1847. The study further reveals that the induced transverse velocity fields modulate near-wall turbulence structures, contributing to reduced turbulent kinetic energy and Reynolds stresses. This method provides a mechanically simpler and potentially scalable alternative for active flow control, with the potential to provide new insights into the mechanisms of transverse forcing in turbulent boundary layers.

A proven concept of the past that yields great potential in the reduction of shipping GHG's, is wind assisted ship propulsion or for short WASP. In particular, the Magnus effect driven Flettner Rotor, a rotating cylinder concept placed upright on the deck of a ship, is on the forefront of promising candidates, mainly due to its substantial aerodynamic force producing capabilities and its relatively low power consumption compared to more conventional ship propulsion.

To accommodate aerodynamic rotating cylinder research at the Low Speed Laboratory of the TU Delft, this study presents the design, realisation and characterisation of an experimental two-dimensional rotating cylinder setup for the SLT wind tunnel facility. Additionally, the effects of testing at a high blockage ratio were investigated for this initial study using the new setup. As a result, the area model blockage ratio deliberately reached an unconventional high 33%, where an aspect ratio of 4.5 was chosen, comparable to full scale WASP applications. The aerodynamic characteristics for three test cases at subcritical Reynolds numbers equal to 62500, 125000 and 250000 with spin ratio ranges equal to 0 ≤ k ≤ 8, 0 ≤ k ≤ 4 and 0 ≤ k ≤ 2 respectively, were obtained by the utilisation of force measurements, particle image velocimetry flow visualisation and proper orthogonal decomposition techniques.

As with similar rotating cylinder experiments, the magnitudes of the presented force coefficients differ in comparison, owing to the large blockage ratio and variations in experimental setup, however the experimental trends in the force coefficient curves are fully captured up to spin ratios of 2.5. Between 0 ≤ k ≤ 0.8, the aerodynamic characteristics are affected by both the Reynolds number and spin ratio, where differences in boundary layer transition, laminar separation bubble formation, boundary layer lengths and vortex shedding behaviour are the underlying mechanisms causing a change in the obtained force coefficients. Transition effects are stronger at the higher subcritical Reynolds numbers, where reversal of the Magnus effect can occur within this range of spin ratios and Reynolds numbers. Beyond k > 0.8, advancing side boundary layer transition has taken place for all tested Reynolds numbers. The force coefficients and flow field topology showed great similarities between the different Reynolds numbers with a further increase of the spin ratio. Slight variations in turbulent kinetic energy production were observed in the retreating side boundary layer and shear layer behaviour, possibly owing to laminar to turbulent transition. For k > 2.0, the periodic shedding of vortical structures is no longer observed, which coincides with the formation of the highly turbulent rotating boundary layer, settling of the near-wake, and the knee in the force coefficient curves within the range of 2.2 < k < 2.6. For remainder of the spin ratios, k > 2.6, the flow field has completely reversed and is pointed in opposite direction of the incoming free stream, where large negative mean drag coefficients and instantaneous force fluctuations are recorded. This is believed to be the result of unsteady nozzle-model gap interactions with the separated region of the cylinder and is dictated by the large blockage ratio.

The Delft Laminar Hump (DeLaH), discovered by the Aerodynamics Department of the Faculty of Aerospace Engineering at the Delft University of Technology, is a symmetrical smooth hump that is placed at a set distance parallel to the leading edge of the wing that reduces skin friction drag as it attenuates the growth of the crossflow instabilities (CFI). With the hump, transition can be delayed up to 14%. The only requirements are that the hump needs an natural laminar flow (NLF) airfoil and a condition where CFI is the dominant transition mechanism to be effective. Generally, CFI dominates when the wing sweep is greater than 30 − 35 deg, making the vertical stabilizer the best candidate for implementing the hump.

This research implements the Delft Laminar Hump on the vertical stabilizer of subsonic transport aircraft by modeling the effect of the hump as a shift in transition location. By using a Quasi-3D aerodynamic analysis in combination with a transition location database, the effect of the hump on the lift and drag coefficient of the vertical stabilizer is analyzed. The transition location database is constructed by using the external velocity of airfoil sections of the vertical tailplane and the boundary layer solver and stability analysis developed by the Group of Flow Control and Stability within the Delft University of Technology. With this, the N-factor curves along the chord can be calculated. Knowing the respective N-factors of the clean and hump configuration, the associated transition locations can be determined, which are then used to calculate the lift and drag coefficients of the vertical tailplane. To evaluate the aerodynamic effect of the hump on the full aircraft directional and lateral stability a stability analysis based on a method by Fokker / Obert is performed and checked against the CS-25 for Large Aeroplanes regulations by European Aviation Safety Authority (EASA).

It was found that the Delft Laminar Hump (DeLaH) on the vertical stabilizer of a subsonic transport aircraft does not affect the vertical tailplane lift curve slope, thus not affecting the stability of the aircraft. In contrast, the vertical tailplane drag coefficient is reduced by the hump. Retrofitting the hump on Airbus A320 (conventional tail) and the Fokker F-28 Mk1000 (T-tail) results in a reduction of the vertical tailplane drag coefficient of 6.73% and 8.72%, respectively. Translating this vertical tail drag reduction to the full aircraft drag coefficient results in a reduction of 0.17% and 0.34%. To evaluate the effect of the hump on weight and fuel consumption, additional weight and mission analyses are performed. Evaluating the harmonic range, an fuel reduction due to the hump of 0.16% and 0.32% is established, for the Airbus A320 and Fokker F-28 Mk1000, respectively. The aircraft weight is reduced by the same percentage through the fuel reduction, as it is assumed that the added weight due to the hump itself is negligible.

Additionally, two sensitivity analyses were performed, namely, sweep angle variation and surface area scaling to analyze the effectiveness of the hump for different vertical tailplane geometries. The effect of sweep angle on hump effectiveness does not affect the vertical tailplane lift curve slope and thus also not the stability coefficients. For the full aircraft drag coefficient, the hump effectiveness has an exponential relation with sweep angle and is most effective at lower sweep angles. A maximum full aircraft drag reduction was found of 0.41% at 30 deg sweep with an equivalent fuel reduction of 0.39%. Overall it is concluded that lower sweep angles are beneficial as the hump is most effective and results in reduced vertical tailplane weight, less fuel weight as well as an increased stability margin. Analyzing the effect of surface area scaling, the hump has no effect on the vertical tailplane lift curve slope regardless of surface area, again retaining the aircraft’s stability. For the full aircraft drag coefficient, the hump effectiveness increases linearly for increasing surface area up to 0.37% at a surface scaling factor of 1.2 times the original vertical tailplane surface area. In terms of fuel reduction, a maximum value of 0.35% was found. There will be an optimal vertical tailplane surface area, since the hump effectiveness increases for increasing surface area, whilst for the full aircraft drag, vertical tailplane weight, and fuel weight a smaller surface area is preferred. The stability margin becomes the limiting factor as a minimum surface area is required for sufficient stability. Comparing the baseline aircraft, the hump is more effective for the Fokker F-28 Mk1000 over the entire range of scaling factors and sweep angles. This leads to the suspicion that taper- and aspect ratio, and cruise speed play an important role in the effectiveness of the hump, but more research is required.

This research shows that the Delft Laminar Hump has a significant drag-reducing effect on the overall aircraft, whilst not affecting the aircraft’s stability. Even though the fuel savings for an individual aircraft are not very large, on a fleet level this would be significant. The hump can be retrofitted on existing aircraft by gluing it on the outer skin, making it a relatively simple and cheap way to improve efficiency for aircraft manufacturers. Nevertheless, more research is required before implementing the hump on commercial aircraft as it is still unknown whether the hump also works on the suction side of the wing as well as whether the hump causes a shock at cruise Mach. Also, interaction effects with the horizontal stabilizer and fuselage need to be taken into account, to fully quantify the effectiveness of the hump. Dedicated wind tunnel experiments, flight tests, and/or CFD simulations are necessary.

This thesis investigates the impact of additive manufacturing induced surface roughness on the heat transfer mechanisms within gas turbine cooling channels. The study involves the design and utilization of a Particle Image Velocimetry (PIV) experimental rig. The research focuses on understanding how the AM roughness affects the flow behavior and heat transfer in micro-channels. Results demonstrate that flow stagnation points significantly contribute to the heat transfer enhancement. High turbulence regions, especially following large roughness elements show an increased heat transfer. Secondary flow features were identified close to the wall, like high-speed streaks. However, the findings show the necessity to further develop the PIV system for closer wall measurements to capture localized effects due to the roughness.

There is a need to examine the potential of alternative methods such as Variational Multiscale Method (VMM) for the ability to improve the efficiency of combustion simulations. However the computational expense of a LES run of a 3D turbulent non-premixed combustion problem is high due to the large variety of length and time scales in turbulent non-premixed combustion. Secondly, VMM is a sophisticated model and the models used to simulate turbulent non-premixed combustion are complex. And finally, the results of a full 3D turbulent non-premixed combustion simulation are difficult to interpret due to their complexity. Therefore the objectives as follows: Firstly to define a simplified model problem which has the important aspects of turbulent non-premixed combustion for comparing numerical methods. Secondly, to propose and to investigate the potential of a framework based on VMM formulation for non-premixed combustion computations. The results of the simplified model problem in a VMM framework are compared to the results when using a standard Smagorinsky model. Results show that in general the VMM model with the dynamic subscale assumption outperforms the Smagorinsky model when looking at the mean and RMS value of the velocity, mixture fraction and progress variable.

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.

Research into constructing reduced-order models (ROM) to reduce computational cost or to interpret complex datasets from numerical or experimental sources has steadily been gaining popularity over the past few years. Many methods currently exist to construct ROMs. The proper orthogonal decomposition (POD) is one of the most widely adopted methods used for these purposes in the field of fluid dynamics. The POD modes favor high-energy structures due to the formulation of the POD however. This may result in an inaccurate ROM, as it might be possible that dynamically relevant, yet low-energy structures are truncated in favor of higher-energy structures. In recent years, an alternative goal-oriented method has been proposed. This goal-oriented method not only has less limits on the goal function that is optimally represented relative to thePOD, the resulting reconstruction of the flow field is model constrained as well. Both these properties are conjectured to lead to more dynamically relevant modes. This goal-oriented, reduced-order modeling (GOROM) technique has been formalized for the advection, Burger’s and incompressible Navier-Stokes equations in 1D and the Stokes equations in 2D, all with homogeneous boundary conditions. This work extends the formulation to the 3D incompressible Navier-Stokes equations with inhomogeneous boundary conditions and provides a parallel implementation. The method is applied to a DNS dataset of a 2D transitional boundary layer with and without a forward-facing step present and compared to the performance of the POD. It is found that the GOROM modes provide a basis which is numerically more accurate than the POD modes for most cases, yet do not provide additional insight into the flow physics. The results presented nevertheless show promise for further applications to construct more accurate and goal-oriented ROMs with the GOROM technique, especially for problems with nonlinear interactions and nonlinear goal functions.

Future designs of turbofan engines are heading towards High By-Pass Ratios (HBPR) to satisfy the requirements of improving the overall efficiency, but there is an inherent acoustic penalty related to this. An increase in the by-pass ratio imposes structural constraints on the engine leading to a decrease in the axial distance between the rotor and the stator. The turbulent fan wake thus becomes more coherent on the OGVs, resulting in an increased tonal noise production. A test setup which can emulate the noise production mechanisms on a smaller scale and study the noise mitigation process is currently lacking. A characterization experiment of such a simplified test setup to assess its ability to predict the vortex-cascade interaction noise as in a realistic fan-stage is performed in this master thesis. The setup consists of an arrangement of linear cascade of OGVs, and a rod which mimics the rotor of a turbofan engine. Two experimental campaigns are undertaken. The first one consists of a wing (NACA 5406) and a rod mounted on flat plates. The wing mimics the mean aerodynamic loading of the central blade of the cascade. This experiment is performed to study the individual characteristics of the OGV blades, when placed in a cascade. The second setup is that of the Rod Linear Cascade model. Far-field directivity of noise sources are analysed using microphones placed on directivity arcs. Ability to quantify the upstream propagation of noise from the cascades is realized using a beamforming array, which is recessed on the test section wall and facing the pressure side of the OGVs. The flow-field at different locations in the test section is analysed using Hot-Wire Anemometry. The Hot-Wire measurements are also used to supplement the acoustic results. The test section is observed to successfully emulate the vortex-cascade interaction noise and this noise is observed to scale with the sixth power of flow speed. The beamforming array is found to help in improving the SNR of acoustic measurements in the test section and hence allows the quantification of the upstream propagation of vortex-cascade interaction noise.

For many aerospace applications, the dominant airfoil self-noise source is Turbulent Boundary Layer Trailing Edge (TBL-TE) noise. The replacement of solid airfoil trailing edges with permeable materials proved to be an adequate noise reduction mean in previous studies. This thesis project contributes to the development of innovative permeable materials by assessing the feasibility of actively influencing their noise mitigation characteristics. The proposed activation mechanism consists of heating up a polymerically coated, porous trailing edge and thereby influencing the material geometry and seepage fluid properties. Heating of the polymeric coating layer led to slightly decreasing pore diameters due to thermal expansion. The main limitation of actively changing the porous material geometry on a pore-scale level was the restriction of the coating layer thickness. Heating effects on the seepage flow through porous metal foams were analyzed experimentally. Characteristic material parameters, namely flow resistivity, permeability and form coefficient, were measured for varying fluid temperatures and it was shown that for low seepage velocities, the pressure communication across the material was negatively affected upon heating. The feasibility of actively changing far-field noise characteristics was demonstrated based on acoustic measurements in an anechoic wind tunnel. Increasing sound pressure levels were observed for both, coated and uncoated porous trailing edges upon heating. It is concluded that the dominant activation effect was the reduced communication across the porous trailing edge due to an increase in fluid temperature. The expansion of polymeric coating was not sufficient to alter the noise mitigation behavior of the porous trailing edge. Furthermore, no interactions between turbulent boundary layer wall pressure fluctuations and the soft coating layer were observed. However, it was shown that temperature control of the porous material offers a possibility to actively influence flow resistivity without modifying the geometrical structure.

The high efficiency of open rotor propulsion has led to a comeback of propeller propulsion systems in recent years. The main issue preventing wide application of the technology is its high noise level. One of the main contributors to this noise is the pressure fluctuations generated during the interaction between a propeller tip vortex and a downstream surface.

The goal of this research project is to create a better understanding of the vortex- surface interaction, which would allow for more effective noise mitigation strategies to be developed. An experimental approach was adopted to quantify the effect of a number of governing parameters on the pressure fluctuations over the surface induced by the vortex. Furthermore, an attempt was made to develop a conceptual model describing the vortex-surface interaction in order to study the relative contributions of the sub-phenomena of the interaction. The accuracy of the conceptual model was assessed using the data obtained from the experimental campaign.

From the results it was clear that the low pressure vortex core was the primary mechanism that generated fluctuations over the surface. The vortex path was clearly visible and dominant over a large part of the chord. In the slipstream region the wake generated the largest pressure fluctuations. The effect of the wake was contained at the leading edge, up to 4% of the wing chord.

Large discrepancies between the unsteady pressure, both in magnitude and in spatial distribution, were observed. Three contributing factors to the discrepancies were identified, namely the overestimation of the induced angle of attack effect, the use of the Lamb-Oseen vortex model and the method of determining the vortex core radius. These discrepancies implied that the conceptual model could not be used for its intended purposes.

The results of this study show that the advance ratio and incidence angle are the most critical parameters governing the vortex-surface interaction. Although the former is difficult to incorporate into noise mitigation strategies, the latter can be used as a parameter in the initial design phase. Interestingly, the geometry of the airfoil seems to have limited effect on both the magnitude and distribution of the interaction. Although the conceptual model discussed in this report is not accurate, the development of such a simplified model would still contribute a lot to the body of knowledge. It could be used to evaluate the relative contributions of the different vortex-surface interaction sub-phenomena, or to perform a 'parameter sweep' to investigate the effect of the governing parameters on the unsteady pressure distribution and magnitude.

Plasma actuators are fully electronic devices without any moving mechanical parts. Hence they have increased reliability and lower complexity when compared to conventional actuators used for same purposes. Mobile components also contribute to increased vibrations, noise, loss of energy and they require lubrication. Non-thermal plasma actuators are highly efficient as they directly convert electrical into kinetic energy without use of any mechanical parts. Lack of mobile components in plasma actuators contribute to lower mass, meaning lower cost of spacecrafts. Another benefit related to use of these actuators is their short response time, which increases dynamical and control abilities of complete spacecraft. All the benefits just stated make this technology potentially very interesting for space-based applications. Presented thesis represents feasibility study, a first step required for application of plasma actuator on a spacecraft.

This research examined two objectives. First was to study the effect of an altitude on momentum exchange between high velocity external flow and created ionic wind. Second objective was to investigate and design DBD plasma actuator able to withstand orbital thermal loads. This study consists of literature research on two most common plasma actuators; various aspects of launch, orbit and re-entry of spacecrafts; and two analytical models. First model aims in estimating the influence of an altitude on the actuator force creation, which can be used for shock stand-off distance modification, and thus steering applications. Other model aims in estimating actuator thermal loads occurring due to spacecraft orbit in Low Earth Orbit. Results show that altitude has large effect on actuator force production. Analytical model shows an average reduction of 4.51 or 4.44 times per 10 km, depending on actuator’s orientation with respect to external flow. Difficulties relating high velocity flow fields such as molecular dissociation and very high friction heat loads are presented. Results show that DBD actuator system can produced significant perturbations between spacecraft surface and natural shockwave. Made perturbations are in form of a compression wave that interacts with natural shock wave and creates a dis-balance of forces proficient for spacecraft reentry steering. Second analytical model shows that there exist actuator materials which can withstand thermal loads due to spacecraft orbital Sun exposure. This thesis shows a promising first order results needed to qualify and apply DBD actuator on spacecraft. However much research on optimization and design are still required.

An aneurysm refers to the local dilation or ballooning of a blood vessel and if left unchecked, almost every aneurysm continues to grow in size until it eventually ruptures. The abdominal aorta is the most common location for an aneurysm to form and an abdominal aortic aneurysm (AAA) can turn lethal in the event of rupture. Since the current clinical practice to classify AAAs is on the basis of their size (diameter), it was decided to investigate the impact of varying geometric parameters viz., the expansion ratio (ER) and the expansion angle (EA), on the pulsatile flow field in a hypothetical axisymmetric AAA (4A) geometry and in particular, investigate the role turbulence. All the combinations of ER and EA lead to a total of 8 cases. The simulations show that a vortex ring is shed just after peak systole in every case. The vortex ring, at inception, is similar for all cases within visual limits. As the ring travels downstream, an azimuthal instability sets in that grows with time. This is the short wave or the elliptical instability usually associated with a vortex pair perturbed by an infinitesimally small amplitude sinusoidal wave. The number of waves formed along the circumference of the ring as a result of the instability varies from case to case but in general, the number of waves decreases as the ER increases. The growth of this instability, along with interaction of the ring with the remnants of the flow field form the previous cardiac cycle, cause it to break down into smaller vortices thereby generating turbulent fluctuations. Further, subtle differences between various cases regarding the vortex ring breakdown mechanism are observed and discussed. Investigations are also carried out to identify recirculating or dead flow regions as in the context of an AAA, such a region indicates the presence an intraluminal thrombus (ILT) which is a commonly found feature in AAAs. An ILT deprives the surrounding arterial tissue of oxygen thereby weakening the aortic wall and increasing the risk of rupture. From the simulations, it could be seen that the inflection regions of the 4A geometry are most prone to the formation of an ILT which is in agreement with the physically observed locations. Finally, the oscillatory shear index (OSI) is investigated to see the influence of turbulence on a hemodynamic parameter. It was found that for the geometries tested, there is no clear co-relation between the two. From this thesis, numerous conclusions can be drawn. But perhaps the most important conclusion is that the flow field in an AAA is very complex and sensitive to various input parameters such as the ER and EA. Although flow parameters such as the Reynolds and the Womersley number are kept constant across all the cases, existing literature clearly highlights the dependence of the flow field on them. As such, the desire to be able to formulate general co-relation trends between various flow field quantities in AAAs is wishful thinking at best.