Among all the chaos surrounding the impact of general aviation on climate change, there is a small beacon of peace: the sport of gliding. Gliders are nothing new: they were the very first heavier-than-air aircraft capable of transporting people. It should come to no surprise this sport has only taken off in popularity since then. Typically, a glider becomes airborne via winch or aerotow. Out of these, aerotow tends to be preferred due to thermals guiding the glider’s favoured drop-off location. Surprisingly, the increasing awareness of climate change and unpowered nature of the sport have not led to much research into sustainability of the tow craft used for aerotows. As such, the necessity of designing a sustainable glider tow craft is indisputable. The tow craft designed to fulfil this aim is the Motorised Automatic Return Craft Used for Sustainable Towing, or in short, MARCUS-T. Starting from scratch, with only several top-level requirements given, a set of in-depth analyses of the main subsystems will eventually lead to a preliminary design. In the future, this design should be revised for more detailed design.

Small satellites are a suitable alternative for large spacecraft for various space missions due to their flexibility, rapid production times, and cost-effectiveness. The development of micro-propulsion systems helps maximise the potential of miniaturised spacecraft. TU Delft contributes by researching novel techniques and developing its own small satellites. This project continued the development of a vaporising liquid micro-resistojet, constructed with commonly accessible materials and machines. The main objective was to implement a new setup to improve the accuracy of thrust measurements. Unfortunately, the nozzle geometry deformed during thruster reassembly due to a degraded sealing surface. This rendered comparison with previously obtained data meaningless. Furthermore, predictions made with an analytical model differed from the experimental results. With alterations made to both the thruster and the test bench, identifying the source of this discrepancy became complex. Recommendations were made to improve both the analytical model and the measurement accuracy of the setup.

This study investigates the effect on the principal aerodynamic forces caused by home-built under-wing attachments excluding possible consequences on the aircraft’s control and manoeuvrability. Many other types of wing attachment have already been used and investigated of time. However, none of these attachments match exactly the attachment under investigation in this study.
For aerial display purposes, one of the team members of the Whiskey Formation Team created an under-wing firework launch platform attached to the tie-down points of the aircraft. These tie-down points are fixed to the quarter chord wing spar at about 90% of the half wing span near the wingtip. The launch platforms themselves protrude about a quarter of their own length in front of the wing’s leading edge. The remaining part is attached flush with the wing’s lower skin.
After installing the attachments on the aircraft’s wing, the creator took it upon himself to do some trial high speed taxi runs followed by a standard take-off. A circuit around the field and a full stop landing completed his trial run during which no adverse effects on the aircraft handling or flight safety could be detected.
Next, on the first trial flight as a formation with the launch platforms attached, the team leader claimed he experienced a slight performance increase. Earlier lift-off and sharper turns were among the examples of this performance increase he gave. No claims on the aircraft’s fuel consumption, cruising speed or power-off gliding performance were made. This was not investigated and is deemed not to be relevant to the aerial display.
This report investigates these claims by modelling the aircraft for computational flow simulations, with and without the platforms attached, for both take-off and high G-load turns. For the take-off condition a runway surface is modelled at a wheels-height distance from the aircraft’s wing. Realistic speeds of 50 and 60 knots respectively are chosen together with 10 degrees of geometrical pitch.
The result of these simulations was then validated in real life circumstances to eliminate possible adverse results due to modelling error of the real life aircraft. A number of comparative test flights, 5 in total with 6 take-offs per flight, were performed registering weight and take-off speed for a trimmed stick-free configuration. The results obtained from these test flights are likely to confirm the results obtained from the simulations. For the simulations, a larger lift and lower drag force are noted for the configuration with the launch platform attached for The steep turn case. For the take-off case a drag increase was noted. In the real life test this could be confirmed by the almost identical take-off velocities even though this falls well within the error margin of the experimental setup. The gain in lift, most pronounced for the steep turn case, comes from the blocking of the spanwise flow by the launch platforms hereby likely reducing the lift-induced drag caused by the wing tip vortex. The total drag of the modelled wing is reduced for the steep turn case even though the launch platforms contribute to the friction drag. Moreover, the model’s lift-to-drag ratio, a common measure of aircraft’s performance, is increased.
It can therefore be concluded that for the investigated cases indeed a small effect of the launch platforms on the aircraft’s performance can be found though it is deemed unlikely that this difference is be noticeable. Real life tests only show little promise since the error margin occludes the effects on performance. For flight display purposes a different meaning of the word ’performance’ is meant. The lift force is considered much more important than the drag force, even though the drag force is reduced too. Therefore it can be concluded that the installation of the launch platforms does not adversely affect the aircraft’s performance parameters such as lift and drag. Moreover, it is plausible that they do improve the aircraft’s performance parameters and do improve the performance of the aircraft for aerial display purposes.

The use of plasma actuators for active flow control presents an interesting option for the design of an aircraft. Plasma-actuated systems boast the opportunities of having facilitated maintenance, increased operational efficiency, and a decreased noise profile, with minimal impact on general performance. Yet, due to the novelty of the technology, the exact implementation of a 100% plasma-controlled drone is yet to be done. Group 16 was tasked with designing an unmanned aerial vehicle (UAV) which omits conventional mechanically-actuated control surfaces, and instead uses plasma actuators. This report explores the various technical and non-technical design facets of the Plasma-actuated Unmanned Light Surveillance and Eco-friendly Drone (PULSE Drone). Due to the inherent novelty in the design stages, this paper aims to explain the current PULSE Drone design and its approach such that it becomes a benchmark for future development. From the depth of explanations, this paper becomes a relevant source for future projects aiming to incorporate plasma actuators.

With climate change is an undeniable reality, hydrogen based energy solutions will become increasingly important for specific aspects of economies. To meet emission targets, Solid Oxide Electrolysers (SOECs) and Solid Oxide Fuel Cells (SOFCs) will become an increasingly useful part of global energy infrastructure. Identifying the economically advantageous areas where SOECs and SOFCs can be deployed inside a complex economic and energy systems requires the creation of numerical models which are capable of analysing economic and energy markets as well as models which replicate the characteristics of SOEC/SOFC operations. This thesis describes the fundamental physical principals used to create numerical models which describe the steady state behavior of solid oxide cells and
SOEC systems and discusses the results for various operation strategies of SOEC systems.

When searching for life we tend to imagine faraway exoplanets, and rarely do we think of our
own solar system. Although a lot of focus is put on Mars, Venus, our closest neighbor, could
currently host life. Temperatures of 475 ℃ and pressures 95 times what we experience on
Earth don’t offer the best conditions for life on the surface. However at high altitudes the
temperature and pressure drop until, in the area between 50 and 70 km, they approach those
of Earth. Here, amongst a thick deck of sulfuric acid clouds, it is speculated that life could
exist, perhaps in the form of bacteria living in suspended water bubbles.
The purpose of our mission is to design a remote sensing platform to perform in-situ
measurements on the atmosphere and soil composition of Venus, in an effort to learn more
about the planet while looking for specific biomarkers that could be attributed to life.

Solutions for reducing greenhouse gas emissions are paramount under the current environmental circumstances. With methane and carbon dioxide being the most critical emission gasses, SigmaSat set out to find a way to reduce these emissions and simultaneously fulfill its scientific mission. While executing the scientific mission of designing a small satellite mission to demonstrate the latest advances in artificial intelligence, SigmaSat managed to devise a design that allows players in the energy production industry (such as refineries) to drastically reduce their methane and CO2 emissions.

Converting existing aviation internal combustion engines to run on hydrogen could be a way to make aviation more sustainable. This work is in preparation to actual engine tests, and the goal is therefore to investigate the potential challenges of such a conversion and to assess where the opportunities lie in terms of performance optimization. This is done using a CFD model of an internal combustion engine in ANSYS Forte. Using this model, a form of optimized hydrogen injection is found that increases the power output, while decreasing the NOx emissions. Safe engine operation during testing can be ensured by monitoring the pressure inside the combustion chamber and starting off with low equivalence ratios and retarded spark timings. Further recommendations are made, such as installing injectors in the combustion chamber to facilitate direct injection and replacing the spark plugs for cold-rated ones.

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.

The aviation industry has set ambitious emission reduction targets for 2050. Following the success of electrification in the automotive industry, the potential of hybrid electric propulsion (HEP) has been the subject of many scientific projects. The energy and power of the electric systems that support the aircraft propulsion system are typically applied during the energy-intensive take-off and climb phases of the flight. This results in a considerable mass penalty for the remainder of the mission. An all-electric unmanned aerial vehicle, which decouples from the main aircraft at the cruise altitude and returns the electric propulsive system to the departure airport, could thus allow to limit the environmental footprint of a particular flight. The present study has quantified the potential fuel, energy and emission reductions of an Airbus A320neo assisted by such an auxiliary aircraft. This potential is also compared to a parallel-hybrid A320neo entering into service in 2035. The required modelling and mission evaluations have been performed with the MATLAB-based integrated simulation framework ‘Mission, Aircraft and System Simulation’ (MASS). The mission, aircraft and engine models of this tool, developed at the Royal Netherlands Aerospace Centre (NLR) throughout multiple HEP studies, have been modified. Besides, constraints for engine temperatures and shaft speeds have been applied. For the 3,500 nautical miles (6,482 kilometres) design mission, fuel and energy reductions of about four and two and a half per cent are found respectively. For the 850 nautical miles (1,574.2 kilometres) standard mission, a fuel reduction of about five per cent can be achieved. The overall energy demand increases slightly in comparison with the reference Airbus A320neo. No feasible parallel hybrid-electric design point was identified for these two mission profiles. In order to account for the uncertainty inherent to a preliminary design stage, the impact of multiple model parameter variations is assessed in a design study. In the vast majority of the investigated scenarios, single digit fuel and emissions savings can still be realised with this innovative aircraft architecture. A more detailed aerodynamic analysis and further refinement of the propulsive system model are recommended subjects for subsequent research.

The aviation sector accounted for 3.4 to 4.0% of the total anthropogenic radiative forcing in 2015. The sectors’growth is several percentage points larger than the annual fuel efficiency gain, resulting in an increasing climatechange impact. The world formulated goals for 2050 to limit anthropogenic climate change, the aviation sectorrequires a mix of fuel consumption reducing technology, increased operational efficiency and carbon offsettingto contribute to the goals.The Aviation Integrated Model (AIM) developed at the University of Cambridge resolves the impact of tech-nological and operational choices, but requires detailed knowledge of the technological development. This workis directed to retain the air traffic and fleet resolution model of AIM, while utilising a different performancemodel. Aircraft performance is modelled by technology trends to assess a range of technology options (such ashybrid electric flight or different aerodynamic configurations) at limited required knowledge while retaining themodel’s physical basis.The model uses the Breguet range equation and fuel fractions updated with Lissys’ Piano-X data to resolveaircraft fuel consumption with limited modelling error. The fuel and carbon dioxide emissions have an averagenormalised root-mean-square error of three percent compared to AIM, while nitrous oxide emissions incur a1.5% error.A case study to demonstrate the capability of the model has been performed to investigate the possibilities forhybrid electric flight serving a significant market share before 2050. The current trends in annual fuel efficiencyincrease and increase in battery specific energy do not result in a significant market share or reduction of aviationemissions for hybrid electric aircraft before 2050.To achieve a 10% decrease in carbon dioxide emissions, the annual increase of the aircraft fuel efficiency hasto be increased from 1.1% per year to 1.5% and the annual battery specific energy increase is required to goup to 5.67%. The required increase in both battery specific energy and the aircraft performance parameters issignificant, which makes a substantial reduction of emissions of the aviation sector by hybrid electric aircraftbefore 2050 improbable.

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.

The world is faced with two concurrent trends. The first is the growing demand for air travel in the long term. The second is the need to cut aircraft emissions due to environmental goals and ambitions set by authorities. Fully-electric aircraft are currently infeasible due to the limited battery specific energy and power. Hybrid-electric propulsion (HEP) technology offers an intermediate solution. HEP can work in tandem with new aircraft designs, which potentially leads to further reductions in fuel and energy consumption. This study aims to quantify those reductions. A parallel HEP aircraft mission simulation framework 'MASS' has been developed at Netherlands Aerospace Centre. Earlier studies with MASS have pointed out that HEP retrofit aircraft have the potential to reduce fuel consumption. The expectation is that even more fuel can be saved by varying the HEP parameters and airframe geometry together. Hence, this study focuses on extending MASS to incorporate new aircraft designs. The extension is made in Matlab and involves adding the capability to calculate aircraft level aerodynamics and weight based on the geometry. A tool called 'OpenVSP' is used to generate the aircraft geometry. Then, the aircraft drag polar is constructed using a combination of the aerodynamic solver 'VSPAero' and handbook methods. Maximum deviation between predicted and actual drag for the relevant lift coefficients is 7.3%. The weight is calculated using Class-II estimation methods. Two separate sets of formulae are used: One is meant for the tube-and-wing (TAW) configuration whereas the other is dedicated to blended wing-body (BWB). The predicted operational empty mass deviates less than 1% from the reference for TAW aircraft, and 2.3% for BWB. HEP-related research is currently focused on large passenger aircraft. In this study an A320neo aircraft is used as reference aircraft, which is parallel HEP retrofitted. It is subjected to hybrid-electric powertrain and geometry variations in order to quantify the potential fuel burn reductions. During the variation studies, the same mission of 800 nm and 150 pax is simulated. The selected powertrain design variables are the engine core scaling and the power split. The latter is the electric power provided by the batteries to the engine, divided by the total amount of power. The idea of using the power split is to support the downscaled engine core during high thrust phases, which would otherwise not have enough power. A design of experiments is conducted prior to the unconstrained optimization for minimal fuel burn. The optimization is performed directly using Matlab function 'fmincon'. The powertrain optimum shows a 3.1% fuel burn reduction. The geometry variation is done in two ways: one involves planform changes and the other a configuration change to BWB. The optimization with powertrain and planform variables is done via a surrogate model obtained using an artificial neural network. The optimum shows a 16.9% reduction in fuel burn, where the planform contributes more to the reduction than the powertrain. For the BWB fuel burn quantification the same process is applied. The powertrain optimization shows a 3% fuel burn reduction with respect to its non-HEP variant. A 13.8% fuel reduction is obtained with powertrain and planform variations. Again, the planform is a larger determinant of fuel burn reduction than the powertrain. Compared with the non-HEP A320neo, a maximum of 19.9% fuel burn reduction can be achieved.