In recent years, the civil aviation sector has increasingly focused on optimizing propulsion designs to improve efficiency and reduce CO₂ emissions. Propellers have regained attention for their superior efficiency, particularly in applications like full-electric aircraft, though high noise emissions remain a significant challenge. This thesis investigates the aerodynamic and acoustic performance of a Boundary Layer Ingesting (BLI) propeller installed at the end of the fuselage of a transport plane, behind the vertical tail. The study focuses on the impact of unsteady blade loading on performance and explores whether increasing blade count or modifying blade sweep can mitigate drawbacks in a BLI configuration.
The development of numerical tools was the first step in this investigation. The aerodynamic tool, an enhanced Unsteady Non-Linear Vortex Lattice Method, accounts for viscosity and flow separation, while the acoustic solver uses Hanson's Helicoidal Surface Theory to model noise emissions from propellers in both isolated and installed conditions. Simulations were conducted on a modified 6-bladed XPROP-S. The velocity profiles at the propeller disk, used as inflow for the installed BLI analysis, were taken from CFD results of the fuselage with the vertical tail installed.
The complex inflow field led to periodic increases in blade loading, especially behind the vertical tail, causing oscillations in thrust and power. Increasing the blade count from two to six resulted in steadier performance and reduced force oscillations. In the installed configuration, the propellers exhibited a thrust-to-power ratio increase of 2% to 5%, with in-plane forces scaling with thrust and remaining within 3% of it. Noise emissions rose significantly after installation, with peak emissions shifting from the rotational plane to the propeller axis, where unsteady loading noise sources became more prominent. The noise level increase reached up to 46 dB during cruise for the 6-bladed baseline. While increasing the blade count reduced noise levels in isolated conditions due to lower propagation efficiency of steady loading noise, in installed conditions, unsteady loading noise sources counterbalanced this, leading to overall increased noise. Maximum noise emissions ranged from 32 dB (isolated, six blades) to 78 dB (installed, six blades). Sweeping the baseline blade backward at the tip and forward near the hub increased loading and thrust by up to 20%, particularly at the tips. In installed conditions, backward sweep reduced overall blade loading changes, while forward sweep increased them. Noise levels varied, with isolated conditions showing a slight increase and installed conditions experiencing up to a 4.7 dB reduction with backward sweep.
The analysis underscores the importance of accounting for unsteady blade loading and how the dominance of unsteady loading noise sources is influenced by loading oscillation amplitude, propeller geometry, and operating conditions.

To make propellers more viable as a greener alternative to jet engines, the issue of propeller noise needs to be addressed. Previous propeller design optimization studies rarely take into account the effects on aerodynamic and aeroacoustic performance due to non-uniform inflow, the type of flow the propeller would experience if it was installed on an aircraft. This study compares propeller designs optimized in uniform flow with propeller designs optimized in an inflow under an angle of attack of 5 degrees, and compares their aerodynamic and aeroacoustic performance when both propellers operate under an angle of attack of 5 degrees. The objective was to investigate whether it makes a difference in terms of performance if the non-uniform inflow is included and accounted for in the optimization routine. It was found that for the non-uniform inflow studied in this work, there was no significant improvement in aerodynamic and aeroacoustic performance when comparing the design optimized in the non-uniform flow with the design optimized in uniform flow.

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.

Propellers have a high propulsive efficiency and can help in the reduction of fuel consumption of aircraft. For propellers which are mounted in front of the wing, the interaction of the slipstream with the lifting surfaces can alter the lift distribution significantly. In the use of modern optimization tools, fast and accurate analysis of the wing aerodynamics is needed, including the influence of the propeller slipstream. Current potential flow based wing analysis tools, such as the lifting-line and vortex lattice method, highly overestimate the lift increase due to the axial induced velocity component in the slipstream, compared to wind tunnel and CFD data. The shortcomings of the existing wing analysis tools are found in the approach of adding the slipstream velocities to the wing induced circulation. At the wing segments that are submerged in the propeller slipstream, it is assumed that the slipstream is infinity large. To account for the effect of the finite slipstream dimensions on the lift distribution, potential flow methods based on the image vortex technique, are used to correct the existing analysis tools. These tools were validated using CFD simulations using the Euler equations. From the validation process, one method was found to produce a good agreement with the CFD data and this can be used to improve the modeling of propeller-wing interactions.

In this thesis project, it is investigated to what extent the energy consumption of an electric aircraft can be improved by using a variable pitch and RPM propeller, including energy recuperation during the descent, with respect to a constant pitch propeller. The largest percentage of total mission energy saved using the variable pitch propeller with respect to the constant pitch propeller is about 4.1% for a cruise distance of 5 km when using two blades and about 3.0% for three blades. This percentage reduces to approximately 0.7%, both for two and three blades, when the cruise distance is 200 km. A two-bladed propeller therefore turns out to be the best option. Regarding the effect of the propeller airfoil camber and thickness with respect to the NACA 4415 airfoil, the NACA 0015 airfoil performs best on a short range mission and the NACA 2415 performs best on a long range mission.

This thesis experimentally assesses the benefit of Boundary Layer Ingestion (BLI) at the unaccelerated condition (thrust equals drag) of an integrated system in an industrial wind tunnel setting. The boundary layer or wake flow of the fuselage is ingested by the propulsor. Previous research shows that ingesting the boundary layer can lead to beneficial savings in terms of power compared to a compared to a non-BLI propulsor. Traditional thrust-bookkeeping methods are not suitable for BLI systems because of the ambiguity between thrust and drag of BLI systems. In order to analyse the performance of such integrated systems the theoretical framework developed by Mark Drela is used: the Power Balance Method (PBM). Furthermore, the classical definition of propulsive efficiency, also named "Froude" efficiency, is not valid for these systems, thus a relatively new propulsive efficiency definition is used. The wind tunnel tests were carried at the Low Speed Wind tunnel (LST) from the German-Dutch Wind tunnels (DNW) located in Marknesse. Four different configurations were tested in the wind tunnel: i) isolated propulsor, ii) isolated propulsor including an upstream strut, iii) BLI configuration and iv) Wake Ingestion (WI) configuration. The isolated propulsor, simply a propulsor in free-stream flow, is used as a reference to compare with the BLI system. The unaccelerated condition at which these systems were tested is at Mach number M = 0:176. The mechanical flow power, jet dissipation and propulsive efficiency are quantified experimentally by using a five-hole probe. Other relevant physical parameters are also measured such revolutions per minute (RPM), thrust and body drag. The designed fuselage is axisymmetric simplifying the measurements and computations. The results from the experiment reveal that BLI systems have power savings around 29% 2:9% when using the electrical power as a reference compared to the non-BLI system. For WI the power savings are slightly lower 24% 2:9%. Moreover, the power savings have also been computed using the mechanical flow power, but a larger variance has been observed in the results from 29:25% up to 56:27% for the BLI case. The flow specially at the outlet plane is shown to be asymmetric between the starboard and port sides. Furthermore, the jet dissipation decreases by 27% compared to a non-BLI system. The slow moving wake flow of the fuselage is re-energized leading to a decrease in jet dissipation. The propulsive efficiency increases between 8% to 10% for both BLI and WI configurations compared to the free-stream propulsor.

Large Eddy Simulations (LES) of a novel type of wing/body junction called the anti-fairing are performed in the current thesis to study the complex turbulent flow physics involved in the junction area and also to obtain a clear understanding of the drag reduction capabilities of the anti-fairing. In regards to that, two separate LES are performed: one for the baseline case with a Rood wing/flat plate combination and another with the Rood wing/anti-fairing combination. A detailed comparative study is performed between the two cases to observe important differences in junction flow characteristics. Both the simulations are performed on a 25 million immersed boundary Cartesian mesh by solving the incompressible Navier-Stokes equations using the in-house finite volume LES solver called INCA. Results from the LES study confirms the existence of the propulsive pressure mechanism of drag reduction for the anti-fairing case, previously proposed by Belligoli et al. However, the results also show that there exists a secondary drag reduction mechanism caused by a combination of increase in approach boundary layer momentum thickness and dampening of the turbulence associated with the horseshoe vortex (HSV) upstream of the wing. This secondary mechanism has been found to be caused by the convex dent present at the start of the anti-fairing geometry. The total drag reduction for the anti-fairing case comes out to be 1.8%. A new parameter called junction drag is defined which accounts for the drag only due to the presence of a junction. The reduction in junction drag obtained for the anti-fairing case is about 6.8%. Apart from the LES analysis, a RANS analysis has also been performed to further investigate the drag reduction capabilities of anti-fairing for different approach boundary layer thicknesses and anti-fairing depths. All the RANS analysis have been performed on a 5 million body-fitted mesh by solving the incompressible Navier-Stokes using the open source finite-volume solver OpenFOAM. Results from the RANS analysis indicate that there exists an optimum depth for the anti-fairing which corresponds to the least drag. Furthermore, it is found that the effect of approach boundary layer thickness is mostly on changing the base drag of the case where no anti-fairing is present, rather than actually affecting the performance of the anti-fairing at different depths.

The longitudinal characteristics of a multi cranked wing in tandem with a canard were investigated on a 2\% scaled generic, modular wind tunnel model in the 2.85m x 2.85m Open Jet Facility (OJF) at Delft University of Technology. The experimental campaign yielded force, moment and surface oil flow data for a moderate angle of attack range of –5 to +22 °, at a free stream velocity of 30 m/s, and at a Reynolds number of 0.5 million based on the wing mean aerodynamic chord. An experimental approach was followed to understand the complex aerodynamic characteristics of the multi cranked wing. The results presented in this thesis provide the low speed trends for a canard and multi cranked wing configuration, with emphasis on the canard and wing interactions. It provides insights on optimal longitudinal and vertical placement of the canard with respect to the multi-cranked wing for maximizing improvements with respect to lift, drag and longitudinal stability characteristics of the configuration. High fidelity aerodynamic database developed as part of this thesis can be used in the design of future fuel-efficient unconventional canard aircraft configuration(s).

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 Flying V is a concept proposed and patented by Benad and Airbus Operations GmbH, which is a flying wing aircraft in the shape of the letter V. Multiple researches already showed promising results, with a 25% increased lift over drag ratio compared to conventional aircraft like the Airbus A350-900. Before research of sub-scale flight testing can be conducted, research into the flight characteristics is necessary. Based on the identification of the problems, the research question for the thesis research is defined as: "What are the flight characteristics of the Flying V sub-scale model at approach speed and high angles of attack?" In total, three different wind tunnel campaigns were conducted at the Open Jet Facility of the Delft University of Technology. During the wind tunnel campaigns, experiments were conducted to acquire data through balance measurements and flow visualisation techniques by means of oil, tufts ans smoke. The untrimmed maximum lift coefficient was estimated to be 1.09 at 41 degrees angle of attack and positive stability was found up to 20 degrees angle of attack. Flow visualisation results showed that large leading edge vortices acted on the surface of the wing, originating from the root towards the crank of the wing. A leading edge separated vortex spread over the wingtip, which originated at the kink of the leading edge. Investigations into the trimmed flight of the aircraft concluded that due to trim limitations the centre of gravity is bound between 1.345 and 1.425 meter behind the nose. The centre of gravity has an optimal location at 1.365 meter behind the nose, providing a static margin of about 9%. With the trimmed lift curve and the optimal centre of gravity location, the flight speed of 35 meter per second can be flown at 3.6 degrees angle of attack. Due to limitation in maximum deflection of the control surfaces, the maximum lift coefficient is estimated at 0.95, at 28.5 degrees angle of attack at a stall speed of 14.8 meter per second. At 20 degrees angle of attack, where the static stability switches from positive to negative, the lift coefficient is estimated at 0.73 at 17.2 meter per second and is defined as the 'safe' stall speed. Taking the FAA and ICAO regulations into account, the approach speed is estimated at 19.2 meter per second, at an angle of attack of 15.9 degrees at a lift coefficient of 0.58.

Increasing pressure of reducing fuel consumption has led to renewed interests in propeller propulsion devices for their high propulsive efficiency. In a propeller propulsion system, the torque applied to the propeller results in angular momentum in propeller slipstream. This angular momentum is not only a form of energy loss that does not contribute to the propulsive performance, but also disturb the wing lift and drag distributions leading to deviated wing performance compared to a clean wing case. A series of design solutions have been introduced for dealing with this term, and current thesis mainly focus on the concept of Swirl Recovery Vane (SRV). SRVs are a set of stationary vanes mounted at downstream of the propeller, aiming at recovering the swirl in the propeller slipstream, which may produce extra thrust without extra power consuming. Previous research mainly focus on design of SRVs for an isolated propeller case. Considering a wing-mounted tractor propeller configuration, with benefit already gained in terms of wing induced drag reduction by wing swirl recovery, the swirl residual in propeller slipstream gives opportunities of introducing SRVs for extra thrust production, such that the system performance can be further improved. Due to the lack of such research, current project is about design of Swirl Recovery Vanes for a maximized system propulsive performance for Wing-Mounted Tractor Propeller Propulsion System at cruise condition. A wing design case is also performed for the same propeller slipstream without installed SRV, by comparison of performance gain of these two cases, the conclusion can be drawn that whether it’s beneficial to introduce SRVs in a typical wing-mounted tractor propeller configuration. A Surrogate-BasedMulti-Fidelity optimization framework is developed for the design purpose. The strategy is to replace the direct optimization of expensive high-fidelity analysis by an iterative re-optimization of a corrected low-fidelitymodel. The low-fidelity method includes a slipstream model in which the velocities are obtained from RANS simulation of an isolated propeller, a lifting line based SRV designmodule and a surface singularity method based wing analysis and design module. The high-fidelity method is the Euler equationbased simulation. Furthermore, an alignment procedure between low- and high-fidelity results is established based on the shape-preserving response prediction(SPRP) algorithm, which assumes the variation of highfidelity results can be predicted by the low-fidelity results. Two main optimization tasks are performed to reach the final objective. The wing twist optimization for minimum induced drag under propeller slipstream indicates that the CDi can reduced up to 5.93%. A series of one-blade SRV design cases are performed at different streamwise and azimuthal positions relative to wing. Results indicate that an upstream installed vane causes an un-expected induced drag increase of a downstream wing through its wake and tip vorticies development. This can be avoided when the vane is moved to downstream of wing. The optimum system performance is obtained when the vane is located at down-going blade side(DBS). By extra thrust produced by vane, a system induced drag reduction of 6.08% drag counts is achieved, which is almost equal to that obtained by wing twist design. Further conclusion is drawn that SRV should be installed as close to wing trailing edge as possible to obtain an maximized system performance in terms of system induced drag reduction.

The near wake of the Ahmed reference body in the presence of steady cross-wind is investigated experimentally by means of large-scale three-dimensional Particle Image Velocimetry (PIV). A novel experimental technique is employed which exploits coaxial imaging and illumination while relying on Helium Filled Soap Bubble (HFSB) seeding as well as on robotic actuation. A 50% replica of the Ahmed body with a 25° slant angle is, therefore, considered in different configurations, among which the headwind condition and two different yaw angles, β1 = +4° and β2 = +8°. A height-based Reynolds number of 1.15 × 10^5 is chosen together with an airspeed of 12 m/s and the Open Jet Facility (OJF) Wind Tunnel of TU Delft is employed to perform the experiment.The relevance of the proposed work lies in the importance of experimental data concerning the Ahmed reference body as well as in the shortage of it with respect to the cross-wind case. The Ahmed body, indeed, once studied to improve the understanding of the flow field generated by road vehicles and to achieve a substantial drag reduction, is nowadays used as Computational Fluid Dynamics (CFD) benchmark and validation tool for automotive applications.To the best of the author’s knowledge, no work is present in the literature which investigates the flow field around the Ahmed body by means of three-dimensional quantitative measurement techniques such as the employed robotic volumetric velocimetry. The proposed study is, therefore, expected to be a significant contribution to the body of knowledge, supplying new valuable data to be used for validation purposes and to inspire future works.

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.

As the complexity of high-tech systems continuously increases, engineers look for possibilities to reduce time and cost of the development of these systems. Architecture-based design enables a front-loaded design process with knowledge reuse. By enabling the automatic synthesis of simulation models, different configurations of an architecture can be realized and simulated efficiently. Current practices are found in the automotive and aerospace industry where architecture-based design is used for the automatic synthesis of multi-physics simulation models. In this way, different architecture options and simulation model variations can be efficiently investigated early in the development process. Multi-body simulations are also frequently used in the conceptual design of complex mechatronic systems. However a suitable methodology to synthesize their simulation models is lacking.

In this thesis, a methodology is developed that makes three necessary improvements to an existing approach that can synthesize only multi-physics models. Firstly, the existing approach assumed that the order of synthesis of the architecture was irrelevant. However, the use of modular and independent simulation models breaks associative links. Therefore, an additional sorting algorithm needs to be introduced during the synthesis of the architecture. An initial successful attempt is made by adding an implementation of Kahn's algorithm. A second addition of the thesis is the definition of a formalized modelling process for modular multi-body simulation models. The formalized process guarantees that the interfaces between simulation models will always work. It also simplifies the modelling approach which leads to the third improvement: clarity of design intent. A procedure was developed and automated that takes away the burden of creating interfaces between simulation models. Consequently, the modeller can focus fully on creating robust models with clear design intent.

The developed methodology is verified with a case study on the conceptual design of a trailing-edge high-lift system. An architecture is defined and parameterized multi-body simulation models are created that realize the components of the architecture. Besides the components that actuate the flap, such as gearboxes, shafts and a motor, four different deployment mechanism types are modelled. The parameterized geometry of the simulation models adapts automatically in order to provide the correct flap trajectory and to form a consistent simulation model. A tool synthesizes automatically all architecture configurations. Interface forces and moments between bodies can be inspected directly and the required actuation torque is found. The sizing of the components is not performed.

The methodology is evaluated for knowledge reuse. For the case-study, this leads from 41\% up to 92\% of reuse of simulation models. However, before these results can be generalized, a trade-off needs to be made from case to case between the time that is saved by the automatic synthesis of simulation models and the time it takes to create the architecture and the compatible simulation models. It can be concluded that the developed methodology enables an architecture-based design approach for complex multi-body simulation models. Furthermore, the methodology is advantageous compared to a traditional design process where all system configurations have to be modelled and analyzed individually.

During aircraft design,multiple tools are utilised to inspect the performance of the configuration. As the design matures, higher fidelity analyses are conducted to predict the flight dynamics of the aircraft. These analyses are conducted by using semi-empirical relations, numerically analysing flow behaviour, conducting wind-tunnel tests and performing scaled test flights. However, Semi-empirical relations might not hold for next generation aircraft and wind tunnel testing and scaled test flights are extensive and are also prone to accuracy issues. A best of both worlds can be found in numerical analysis. However, an increase in flow fidelity modelling comes with an increase in computational cost. Besides, complete analysis of all possible manoeuvres of a design increases the number of computations significantly. Current methods cope with this issue by using flight dynamics models based on so called stability derivatives, instantaneous values which couple flight state parameters to aerodynamic loads to predict aircraft flight dynamics. However, these models do not take into account time dependency. Therefore, these methods do not accurately predict the flight dynamics of agile aircraft, such as unmanned combat aerial vehicles, undergoing rapid manoeuvres where unsteadiness dominates flow behaviour. This conventional reduced-order modelling method, in which samples of the full-order model are taken in the form of stability derivatives, causes design iterations to be analysed inaccurately. The objective of this report is to investigate reduced-order modelling for flight dynamics prediction, thereby comparing conventional techniques to a method which does take into account unsteadiness in flow behaviour.
The method investigated is based on indicial step response functions, which are samples in the form of unsteady aerodynamic flow behaviour functions of the full-order model. The idea is that once these samples are known, any flight manoeuvre can be analysed within minutes. Research found in literature has assessed some of the capabilities and limitations of this method, but not yet applied this to flight dynamics prediction. The research described within this report will address this gap by using two test cases. The first testcase is used to assess the assumptions made in literature, on aerodynamics loads modelling, by applying the method on a two-dimensional airfoil in subsonic flow conditions. It was found that the indicial step response functions are indeed representing the full-order model, thereby taking into account unsteady flow behaviour in aerodynamic loads prediction. In longitudinal motions, the angle of attack and pitch rate effect need to be taken into account to predict lift, drag and pitching moments. Multiple frequencies of the same manoeuvre can be analysed within minutes once the samples are calculated. Results show that the accuracy of the predictions becomes a trade-off issue between samples calculated and accuracy required. The second testcase is used to apply the indicial response functions to flight dynamics prediction of an agile
unmanned bomber aircraft undergoing fast manoeuvres. A longitudinal-directional climbing manoeuvre was calculated by developing a flight dynamics model based on stability derivatives. The flow behaviour encountered during this manoeuvre was analysed to include highly unsteady and non-linear phenomena (e.g. vortices and flow separation) at higher angles of attack. By comparing the results of themethod under investigation to the full-order solutions, it was shown that aerodynamic flight dynamics predictions were accurate in capturing unsteady behaviour and weak non-linear flow behaviour. However, the samples proved to be inaccurate in representing behaviour in highly non-linear regions. Concluding, this means that indicial step response functions provide more accurate flight dynamics predictions than conventional stability derivatives in representing unsteady flow behaviour. The accuracy of the predictions are highly dependent on the samples chosen. Several samples suffice to predict the unsteady behaviour for linear and weak non-linear flow regions of the flightmanoeuvre. If surrogate modelling is applied, the method can become more computational efficient than conducting multiple full-order time-marching numerical calculations. It is recommended that more research is performed on indicial step response functions
in capturing highly non-linear flow behaviour, as the research showed that the size of the samples affects the flow behaviour representation.

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

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

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

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

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