The integration of Generative Artificial Intelligence (GenAI) in engineering presents both opportunities and challenges, particularly in complex, safety-critical systems. This thesis explores two key aspects: (1) developing an AI-powered reverse engineering tool for Knowledge-Based Engineering (KBE) applications and (2) designing a communication framework to enhance interdisciplinary collaboration in AI-driven projects. The aerospace engineering component introduces the REProcess tool, leveraging Large Language Models (LLMs) to extract structured process representations from KBE applications, improving traceability and automation. The science communication component investigates collaboration barriers between AI experts and domain specialists, leading to the development of a structured ideation matrix for GenAI-powered automation systems. Together, these contributions demonstrate how Explainable AI principles can facilitate the adoption of GenAI in engineering workflows while bridging communication gaps. This interdisciplinary approach provides practical methodologies to enhance AI-driven engineering processes, ensuring both technical feasibility and effective human-AI collaboration.

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 role of winglet slots in reducing tip vortices, inspired by bird wingtips. The effects of slotting and wingtip flexibility were analyzed using multiple bio-inspired models tested in a low-speed wind tunnel at Reynolds numbers from 2.9 × 10⁴ to 8.8 × 10⁴. Three experimental techniques were employed: force balance for aerodynamic force estimation, motion tracking of wingtip movement, and stereo PIV analysis to measure cross-flow planes in the winglet wake. Non-intrusive methods were used to estimate drag components based on PIV data and wake integral approaches. The slots significantly reduced induced drag by breaking down the tip vortex into smaller, weaker vortices. This breakdown was visualized using s-PIV results and analyzed through velocity component contours. The streamwise vorticity breakdown was examined in relation to wingtip flexibility, and the vortex cores' location and characteristics were analyzed, along with the phenomenon of vortex wandering.

By the end of the thesis, it was concluded that although slotted winglets reduced induced drag, they led to a significant increase in profile drag, resulting in higher overall drag compared to unslotted winglets. Additionally, the slots caused a loss in lift. The overall aerodynamic performance, measured by the lift-to-drag ratio, was also higher for the
unslotted winglet. Among the slotted winglets, static bending was common, while wingtip vibrations were minimal. The most flexible slotted winglet had the lowest induced drag. However, the results indicate that at very low Reynolds numbers, slotted winglets deteriorate the wing’s aerodynamic performance compared to unslotted ones. Among the slotted winglets, those with intermediate flexibility demonstrated the best aerodynamic performance, highlighting the importance of optimizing wingtip flexibility.

To reduce greenhouse gas emissions in aviation, innovative propulsion systems such as battery-powered electric aviation are essential. These systems are carbon neutral when powered by 100% green electricity. However, widespread adoption requires overcoming challenges including improving energy density, lowering costs, and maintaining safe thermal operation (incl. thermal stability). This thesis addresses the challenge of maintaining battery thermal stability during flight by developing a combined electronic circuit and thermal network model for the NLR-owned Pipistrel Velis Electro aircraft. Using flight test data from this aircraft as a validation source, the model evaluates three battery thermal management strategies: two liquid cooling methods (ribbon and cold plate) and one gas cooling method (air cooling).
The electronic equivalent circuit model, used to simulate voltage characteristics and heat production in a single lithium ion battery cell, requires pulse current characterization tests for accurate parameter estimation.
Extensive testing has been conducted to gather these data. The model achieves accurate voltage modeling accuracy with a low root mean square error Adding more than one RC branch to the circuit did not significantly improve the accuracy of the model.
The electronic equivalent circuit and lumped parameter thermal network models were validated with two flight data sets. Both ribbon and cold plate cooling solutions effectively matched the validation temperatures, performing similarly. In contrast, the air cooling solution was less effective. In case of ribbon cooling, the maximum cell temperature was highly sensitive to its geometric parameters, specifically the angle and height of the ribbon. The sensitivity of the cold plate solution in terms of maximum battery temperature was influenced by the diameter of the cooling channel and the thickness of the plate. The air cooling showed sensitivity in terms of maximum battery temperature relative to the inter-cell gap width. For the ribbon model, varying the number of thermal nodes led to a convergence in the maximum battery temperature as the node count increased.
Using the validated ribbon cooling model, two operational scenarios were analyzed. The first scenario involved charging operations, where simulations closely matched temperature validation data, showing only a minor temperature rise in the battery pack. The second scenario tested cold weather operations with ambient temperatures reduced to approximately 0 ◦C. Here, two simulations were conducted: one with the battery preheated to 20 ◦C and another without preheating. Without preheating, the battery pack’s temperature neared the operational lower limit of 0 ◦C. Preheating prevented reaching this lower limit. It is recommended to preheat the battery pack using an external charger, as using the battery’s own energy for preheating is inefficient.
The thesis was concluded by using the developed modeling approach to size and model a battery pack for the Eviation Alice, a larger aircraft. The approach was successfully scaled to this large use case with a known power profile. Furthermore, due to significant ambient temperature effects and higher operational altitudes compared to the Pipistrel Velis Electro, thermal insulation will be necessary for the battery pack to maintain temperatures above the lower operational limit of 0 ◦C during typical missions.

The global emission of harmful greenhouse gasses has to be reduced. Commercial shipping contributes to these emissions by using fossil fuel as the only energy source to provide propulsion and power auxiliary systems. In order to reduce fuel consumption, much research is focused on increasing the efficiency of the propulsion system and reducing the required power consumption. In recent years an increasing amount of research focuses on utilizing wind energy as an alternative energy source. This is a renewable source of energy which is plentiful at sea. This thesis focuses on the novel concept of using a Vertical Axis Wind Turbine (VAWT) as a means of ship propulsion and power generation. Existing Wind Assisted Ship Propulsion (WASP) systems can only be utilized in conditions when the ship is in transit and when the apparent wind conditions are favorable. The VAWT concept has the ability to generate power in wind conditions where conventional sailing systems are not functional, including when the ship is in port. This system can also be used as a thrusting device when the wind conditions are favorable for sailing systems. Therefore this system can be used more frequently which is beneficial for the overall power savings. The use of a VAWT for WASP in combined operation as a thruster and a power generator does currently not exist. Therefore the design of the rotor has to be investigated and optimized in order to evaluate effective designs for different conditions and to compare this device to existing sailing concepts. The aerodynamic performance of the VAWT will be evaluated using numeric simulation models. This model will be coupled to an optimization algorithm in order to find the optimal design parameters for different wind conditions and modes of operation. This project will result in a better understanding of the working principles of VAWT used for WASP and the potential energy saving of optimized designs for the operational modes as a generator, thruster and combined-operation. There is a large opportunity for capturing wind energy at sea. This project has the potential to add a new method of ship propulsion to reduce shipping emissions and reducing the cost of ship operation. When one or multiple of these systems can be applied on a large share of the worldwide shipping fleet, this research can contribute to the reduction of commercial shipping emissions.

This thesis is performed with the aim to investigate whether split flaps can be a feasible solution for increasing the Flying V directional control power. In order to do so, a set of wind tunnel experiments have been concluded with a 4.6% half wing model in the Open-Jet-Facility at the Delft University of Technology. Here, several split flap geometries have been tested on the model outboard wing section to fully establish their behaviour in terms of yaw and other parameters of interest. Additionally, calculations have been performed to provide an adequate split flap design for certification of the lateral-directional specifications from CS-25.

From the wind tunnel experiments, it was found that split flaps on the outboard wing section can effectively increase the directional control power over angles of attack between 0 and 30 degrees. However, yaw control is lost when over negative angles of attack due to lower wing stall. The maximum decrease in split flap effectiveness over positive angles of attack is found around 17.5 degrees, where a leading edge vortex is present over the outboard wing. This decreases pressure over the upper split flap, resulting in a effectiveness decrease of around 41%. As the effectiveness does not decrease further, split flaps can be continuously effective at higher angles of attack when compared to the winglet rudders. The low pressure on the upper flap also causes large adverse coupled moments in pitch and roll. The effect of split flaps on the yawing moment is found to be linear with deflection angle, where their effect on other aerodynamic properties is found to be more non-linear.

The deflection of outboard split flaps do not have a significant interference effect on rudder yaw power, but can have some interference effects on adjacent main wing control surfaces. Differential deflection between the upper and lower flap has been shown to potentially decrease coupling in pitch and roll. The split flaps can also be globally rotated trailing edge down to mitigate adverse coupled moments while beneficially increasing the total created yaw.

From trim calculations, it was found that a maximum additional yaw coefficient of 2.4113e-3 has to be provided by the split flaps. Designs with a maximum deflection angle above 30 degrees are deemed feasible, as this would maximally require only the replacement of the current CS3 surfaces. When a maximum deflection of 60 degrees is set by the designer, a sub-scale split flap width of 111.87 mm is needed, which translates to a full-scale split flap of 2.431 m. Based on the sensitivity analysis and projected Reynolds effects, this is considered a conservative design for the full-scale Flying V. At 0 degrees angle of attack, the recommended geometry is projected to increase the maximum directional control power of the Flying V by 38.1%. At 27.5 degrees, the maximum directional control power can even be increased by 85.5%. Both increases come at the cost of a significant drag penalty.

A new aircraft configuration, the Delft University Unconventional Configuration (DUUC), was recently invented by the Flight Performance and Propulsion group at TU Delft. As a solution to the concerns regarding the environmental and health impact due to increasing demand of air travel. The DUUC has a characteristic feature which consist of two ducted propellers at the empennage of the aircraft with vertical and horizontal vanes mounted downstream of the duct, spanning the whole diameter. With this newly integrated architecture the DUUC combines thrust with stability and control of the aircraft. Furthermore, with the current trend of increasingly larger bypass ratios on turbofan engines, these will at a certain point not fit below the wings anymore. Therefore, further highlighting the importance of this research than the study on the DUUC alone. From previous research on the DUUC, where the propulsive stabilizer was in a fixed, aft-fuselage position with a fixed duct aspect ratio. It has been observed that the rear center of gravity (CG) position and a large shift in CG due to the propulsive stabilizer, poses a number of challenges. Therefore, research has to be done on the positioning of the propulsive stabilizer. The objective of this research is to study the positioning of the ducted fan propulsive stabilizer, considering its impact on stability and control by development of a method to predict aerodynamic/stability & control derivatives for variable position and duct/fan size. In this study, the aerodynamic analysis tool FlightStream is used to determine the aerodynamic characteristics of the ducted propeller. These results are then used in combination with AVL and an ESDU method to determine the influence of the positioning and duct size on the aircraft stability and control. Furthermore, in order to assess the validity of FlightStream, a validation study has been performed as well and compared to analytical methods which have been used on earlier studies on the DUUC.

Optimal design of a lift dumping assembly (LDA) has been found utilizing a gapped spoiler design that consists of two separate surfaces that are placed at 0.06b-0.17b and 0.28b-0.038b spanwise locations. The gapped design allows the cross flow amplified at the hingeline of the inboard spoiler to escape through the middle gap increasing the lift dumping per area efficiency of the design by 44% while only reducing lift dumping by 4.6% compared to the full spoiler span configuration. LDA spoilers reach a projection height of 8.7% local chord and the hingeline is placed at 80% local chord. With this combination of projection height and chordwise location, there are no occurrences of adverse lift phenomena. Other tested spoilers that are either smaller in chord length, deflection angle or placed such that it is submerged in the boundary layer have shown to increase the lift coefficient at angles of attack relevant to touchdown, 17 degrees.

Control surfaces are integrated within the LDA given the spoilers do not heavily influence the authority of the control surfaces. The ground roll performance of the Flying without the LDA is already comparable to a B747-400 using spoilers. At the point of operation where ground roll is maximized, namely forward center of gravity and maximum landing weight, the Flying V can come to a stop from touchdown in 1300 m and 28.4. Introducing the LDA to the Flying V is less effective than introducing spoilers to a B747-400 as the Flying V does not have any lift from flaps that the LDA can spoil. Nonetheless at typical runway conditions LDA reduces the total ground roll by 95 m. The significance of dumping lift is amplified for adverse runway conditions as 130 m in wet runways and 500 m in icy conditions can be gained using the LDA.

At 17 degrees touchdown angle of attack the LDA only uses the spoiler to create a slight nose down moment of -0.0069 to help derotate the aircraft. Furthermore the LDA can be directly deployed at touchdown to decrease the lift coefficient by 0.057 which is approximately equal to decreasing the angle of attack by 1.5 degrees on the clean aircraft reducing the possibility of a bounce-off. At braking attitude of -2.5 degrees angle of attack, control surfaces join the spoiler in lift dumping. In this case the LDA reduces the lift coefficient by 0.21 and increases the drag coefficient by 0.028.

As part of a project to modify one of the engines of a Cessna Skymaster to operate on hydrogen, a test set-up is designed. A crucial part of this design is the Safety Analysis, for which a risk analysis must be performed. In this thesis, an in depth analysis of the probability of one of the risks involved in engine testing with hydrogen is conducted: the risk of catastrophic engine failure as a result of an increase in pressure and/or temperature inside the cylinders. It is found, that, although a small increase in peak temperature and a significant increase in peak power is found for hydrogen compared to Avgas, the risk of catastrophic failure is probably low. A final numerical simulation is proposed to obtain a definitive answer, and experimental validation experiments are proposed to increase the accuracy of the risk analysis.

A collaborative group comprising of the TU Delft, Deltion College and Dutch Electric Aviation Centre aim to develop a flying test bed for sustainable aviation. The collaborative group has acquired a Cessna 337F Skymaster with the intent to replace one of the engines with a retrofit hydrogen internal combustion engine. An ancillary engine will serve as test unit prior to modifying one of the engines of the Cessna 337F Skymaster. Experiments will be performed in a 20-feet container. Hydrogen may leak from a fuel pipe or through a fuel injector from the engine, consequently hydrogen may accumulate in the enclosed testing environment and a detonation occurs. CFD models of the test cell are simulated in ANSYS fluent. One may conclude that heat and hydrogen can be disposed of well, providing an optimal placement of the ventilation system and test cell elements.

This work is a part of the Automated Rotor Blade Inspection (ARBI) project. ARBI aims to improve the rotorcraft RTB process with novel rotor blade measurement equipment such as 3D scanning, thermography and shearography. The research aims to quantify the effects of changes in blade properties on rotor vibration levels for CH-47 Chinook. This is done in order to create a rough ranking of which parameters have the highest impact on vibrations. This was performed by changing the blade properties on a single blade in the rotorcraft simulation tool FLIGHTLAB, trimming to a desired flight condition and collecting acceleration measurements. In FLIGHTLAB a novel inflow solver, the Viscous Vortex Particle Method (VPM) was applied, since the developers specified it was well suited for rotorcraft with rotor-on-rotor aerodynamic interaction. The results of this research in future hopefully will be able to help identify blades with poor RTB potential and help with suggesting blade sets with higher RTB success chance based on measurements from ARBI.

To reduce the environmental impact of aviation, new aircraft configurations are investigated. One of these researches is the PARSIFAL project, that investigates an efficient box-wing aircraft for passenger transport. In previous work, an engine sizing tool is developed to design the engines of the PARSIFAL aircraft. The tool is developed as part of the Multi-Model Generator (MMG), a knowledge based engineering application developed in the Flight Performance section. This work also included an engine integration study, which showed that a fuselage mounting is preferred over a wing mounting. However, the integration study lacked the desired design sensitivity for small design changes, e.g. engine diameter and length. Because of this lack of design sensitivity, the question whether the engine installation following from this work is ideal remains open. Therefore, in this thesis a best practice for engine integration studies during the preliminary design phase is investigated. The found best practice is then used to re-asses the engine installation by aligning the engines with the local flow. The best practice is also employed to asses under wing engine installations, to establish the potential use of the practice for this type of installations. Before the trade-off study for a best practice could be performed, changes to the MMG had to be made. Three primitives, the building blocks that create an aircraft geometry, had to be adapted. A new wing primitive, through flow nacelle (TFN) geometry, and a pylon primitive are introduced. In search of a best practice, textbook analyses, 3d panel solvers, and combinations of the two are compared. To fully assess an engine integration, information on lift, drag, and pitching moment is required. Since textbook methods only provide information on drag, these are not satisfactory on their own. Nonetheless, these methods can be used to predict the drag addition of a pylon, when only airframe and nacelle are simulated. In two panel solvers, FlightStream and VSAERO, two analysis practices are tested. The first being a simulation including airframe, nacelle, and pylon and the second being a simulation including airframe and nacelle complemented with a textbook correction for the pylon drag. Following a trade-off, using speed, accuracy, ease-of-use, and design sensitivity as criteria, a simulation with FlightStream of airframe and nacelle with a textbook addition of pylon drag proved to be the best analysis practice. This best practice is employed to re-asses the engine installation of the PARSIFAL aircraft. While the location is fixed, since it aligns with the structure inside the airframe, the orientation is changed to align with the local flow. This alignment results in a 3% decrease of aerodynamic efficiency loss and a 17% reduction of pitching moment increase, with respect to the original engine orientation. In a second study, the potential of the best practice to analyse under wing engine installations is tested by changing the location of an engine under the rear wing of the PARSIFAL box-wing. These results show that with proper placement, the same aerodynamic efficiency loss can be obtained as for fuselage installations.

The current focus in the aviation industry for more sustainable designs, could mean the revive of propeller propulsion, due to their relative high propulsion efficiency compared to jets. In addition, the application of wingtip-mounted propellers installed in tractor configuration can be used as tip-vortex attenuating devices, reducing the wing induced drag. The wingtip-mounted propeller configuration is believed to offer a significant aircraft performance benefit from an aerodynamic perspective. So far, studies on wingtip-mounted propellers mainly concentrated on the aerodynamic interaction effects, disregarding the integration with the airframe and wing-structural mass. This thesis presents a methodology to integrate aerodynamic, aero-propulsive, and aero-structural effects of tip-mounted propellers in the context of a typical turboprop featuring hybrid-electric propulsion. The developed methodology is used to assess the effect of wingtip-mounted propellers on both wing and aircraft level, providing new insights in the potential of the configuration.

This research has been performed with the purpose of understanding the cause of the unstable pitch break experienced by the Flying V at a lift coefficient of CL = 0.95. Without understanding and possibly elimination of the pitch break, the flying V’s usable maximum lift coefficient is lower and the pitching moment behavior and gust response unpredictable. Three possible solutions are investigated empirically, with the ultimate goal of eliminating the unstable pitch break. First, the application of a trip strip to the suction and pressure side of the wing is investigated. Second and third, the implementation of vortilons and fences is investigated. The experiments are performed in the TU Delft’s Open Jet Facility using a 4.6% scaled half-span model of the Flying V. The wind tunnel experiments are evaluated using force and moment data and flow visualization using tufts. The application of trip strips results in an increase in pitching moment coefficient up to a lift coefficient of CL = 0.65 and introduces the unstable pitch break at CL = 0.80 compared to CL = 0.95 for the clean wing. Flow visualizations showed an improvement in the flow over the outboard wing when the trip strips were applied. Combining both observations, it is recommended to apply the trip strips on the outboard wing of the Flying V scaled flight testing model only. Vortilons have shown to have no effect on the unstable pitch break. The combination of a vortilon and the trip strips could decrease the pitching moment coefficient of the flying V up to a lift coefficient of CL = 0.70, depending on the spanwise location of the vortilon. Placing a fence could result in a favorable pitching moment coefficient characteristic, depending on its spanwise location and size. A fence placed at the spanwise location of the leading edge kink results in a decrease of pitching moment coefficient between a lift coefficient of CL = 0.30 and 1.1. Furthermore, a fence located at the leading edge kink and spanning the entire suction surface chord postponed the unstable pitch break experienced by the Flying V. Therefore, it is recommended to install that fence on the Flying V scaled flight testing model.

An experimental investigation was done regarding the propulsion system and airframe integration issues associated with the novel Flying V configuration. The experimental investigations involved a series of wind tunnel tests with a 4.6% scale half model, conducted in the TU Delft OJF low speed wind tunnel. Balance measurements allowed significant interference effects between the wing and engine to be identified. A nacelle mounted total pressure rake enabled the measurement of engine inlet total pressure, both in through-flow nacelle and powered conditions. At landing and take-off velocity (20m/s), an interference drag penalty is observed over the full range of positive incidence angles above 5 degrees, with a maximum of 60 counts (16.5% of isolated wing drag) at an incidence angle of 10 degrees. At incidence angles lower than 5∘ engine operation is somewhat beneficial, with a maximum contribution to thrust due to interference of approximately 20 counts. Engine inflow is shown to be distorted by the presence of the wing, which could contribute to the observed interference drag. Distortion DC(60) values higher than 0.2 are measured in off design conditions, while measurements at cruise condition show a DC(60) value of 0.08. However, distinguishing between loss of engine thrust and increase in airframe drag is not possible with the used setup. Increased suction around the engine intake contributes to added lift and nose down pitching moment at high thrust setpoints and incidence angles between 5 degrees and 12.5 degrees. An increase in nose up pitching moment is observed from 12.5 degrees to 22.5 degrees. To put the measured interference in perspective, the direct effects of the quantified interference on trimmed flight conditions are demonstrated. An increased power requirement of up to +8% is observed in trimmed flight conditions between 22m/s and 31m/s, while a maximum reduction of -11% in power requirement is predicted for velocities higher than 32m/s. The change in required control surface deflection for trim due to engine interference is evaluated to be less than +/ − 2.5 degrees for trimmed level flight. A trimmable flight envelope is demonstrated for maximum climb, level flight and optimum glide with control surface deflections in the range of −10 to 10 degrees.

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.

Because of the expected return of civil supersonic flight, research is required to update supersonic noise regulations to allow public acceptance of civil supersonic flights. To advise ICAO on whether future supersonic aircraft will be able to comply with current regulations this thesis predicts the noise production of future supersonic transport aircraft. For this a low-fidelity aircraft design program was created. This program was validated successfully and based on five produced aircraft designs a noise analysis was performed. This analysis shows that supersonic aircraft will be unlikely to meet the subsonic airport noise regulations. The reason for this is that optimisation for supersonic flight inevitably results in changes that produce more noise compared to subsonic aircraft. The sonic boom noise level of future supersonic transport aircraft can be greatly reduced compared to the levels of Concorde by optimising the aircraft’s shape. However, this will result in increased drag, fuel burn and weight. Therefore it is unlikely that near future supersonic transport aircraft will be acceptable in terms of 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.

The second-generation of supersonic civil transport has to match ambitious targets in terms of noise reduction and efficiency to become economically and environmentally viable. High-fidelity numerical optimization offers a powerful approach to address the complex trade-offs intrinsic to this novel configuration. Past and current research however, despite proving the potential of such design strategy, lacks in deeper insight on final layouts and optimization workflow challenges. Stemming from the necessity to quantify and exploit the potential of modern design tools applied to supersonic aircraft design, this work partially fills the gap in previous research by investigating RANS-based aerodynamic
optimization for both supersonic, transonic and subsonic conditions. The investigation is carried out with the state-of-the-art, gradient-based MDO framework \textit{MACH}, developed at University of Michigan's MDO Lab - which hosted the author for the 14-month research stint. Details of the tool and a brief overview of supersonic aircraft design and modern aerodynamic optimization strategies are reported in the first part of this manuscript.
After circumscribing the research niche, I perform single and multi-point optimization to minimize the drag over an ideal supersonic aircraft flight envelope and assess the influence of physical and numerical parameters on optimization accuracy and reliability. Leading and trailing edge morphing capabilities are introduced to improve the efficiency at transonic and subsonic flight speed by relaxing the trade-offs on clean shape optimization. Benefits in terms of drag reduction are quantified and benchmarked with fixed-edges results. It is observed how the optimized airfoils outperform baseline reference shapes from a minimum of 4\% up to 86\% for different design cases and flight
conditions. The study is then extended to the optimization of a planar, low-aspect-ratio, and low-sweep wing, using the same schematic approach of 2D analysis. I investigate the influence of wing twist alone and twist and shape on cruise performance, obtaining a drag reduction of 6\% and 25\% respectively as the optimizer copes with both viscosity and compressibility effects over the wing. Results for 3D multi-point optimization suggest that the proposed strategy enables a fast and effective design of highly-efficient wings, with drag reduction ranging from a minimum of 24\% up to 74\% for cruise at different speeds and altitudes, including edge deflection. Ultimately, this work provides an extensive and, to the best of author knowledge, unprecedented insight on the optimal design solutions for this specific aircraft configuration and the challenges of the optimization framework. The benefits of RANS-based aerodynamic shape optimization to capture non-intuitive design trade-offs and offer deeper physical insight are ultimately discussed and quantified. Given the promising results in terms of performance improvements and design efficiency, it is hoped that this work will foster the implementation of this method for more comprehensive full-configuration, multidisciplinary supersonic aircraft optimization studies.