Long haul flights account for over half of yearly aviation carbon emissions, while only taking up 6% of the total flights. To reduce the emissions of those flights, hydrogen combustion blended wing body (BWB) aircraft are a promising new technology. A conceptual design tool for hydrogen BWB aircraft is created to evaluate the effects of the cabin design and hydrogen tank placement on the aircraft level performance of the concepts. It is found that for a 350 passenger BWB concept a narrow cabin with 45% of the fuel placed next to the cabin results in the lowest fuel burn. For a 250 passenger BWB concept a narrow cabin with 40% of the fuel next to the cabin offers the lowest fuel burn.  The spread in performance is lower for the 250 passenger concepts than for the 350 passenger concepts, indicating the internal layout has a larger effect on the performance of the 350 passenger concepts. Optimising cruise conditions resulted in high optimal altitudes due to the low wing loading of LH2 BWBs, however limitations of the model with respect to the cruise altitude reduce the certainty of conclusions drawn. Conclusions drawn with respect to the sensitivity of the performance to design parameters are limited by uncertainty in the wave drag estimation of the concepts.

This thesis investigates the aerodynamic performance of various Distributed Electric Propulsion (DEP) configurations using low-fidelity modeling tools. The study is motivated by the growing interest in sustainable aviation and the potential of DEP systems to offer improved aerodynamic efficiency, reduced noise, and greater design flexibility. Conducted in collaboration with Pipistrel Vertical Solutions, the research focuses on evaluating and optimizing different DEP layouts that combine minimum induced loss (MIL) propellers, dedicated for efficient cruise, and lift-augmenting (LA) propellers, designed to enhance lift during take-off and landing.

A key objective of the work is to develop a computationally efficient optimization framework suitable for early-stage design, capable of assessing multiple DEP configurations. To that end, the study integrates propeller design methodologies, slipstream modeling, and a DEP-specific lifting line solver into a single low-fidelity analysis tool. This framework is then coupled with multi-objective optimization algorithms, such as NSGA-II and SMPSO, to explore a wide design space and identify configurations that minimize take-off distance and cruise power requirements.

The thesis compares three main DEP arrangements, varying the number, size, and spanwise placement of MIL and LA propellers. The results show how propeller positioning and interaction effects influence overall aerodynamic performance, and they highlight the value and limitations of low-fidelity models in capturing these phenomena. The insights gained provide a foundation for future development of DEP systems and for refining low-fidelity tools for preliminary aircraft design.

The search for extra-terrestrial life has long fascinated humanity, with Enceladus, one of Saturn's moons, emerging as a prime candidate due to hydrothermal activity and methane detected in its South Pole plumes by the Cassini mission. The ELMO orbiter is designed to explore Enceladus, searching for biosignatures, mapping its surface, and relaying data from two hopper vehicles deployed on the Saturnian moon. Carrying five scientific payload instruments, the orbiter must meet strict constraints, including a total mission cost of $750 million USD and compatibility with the Ariane 64 launcher. The spacecraft provides 6000 m/s Delta-V through a two-stage design: a main orbiter (3400 m/s) and a kick stage (2600 m/s). It also features an 85 m^2 solar array for power generation and a dual-band communication system to handle high data acquisition requirements (15% of operations time).

The lightweight truss structure, made of low-density composites, incorporates MLI for thermal regulation and multiple radiation-shielded electronics vaults. The fixed high-gain antenna ensures efficient Earth communication at Ka, X, and S bands. Weighing 13400 kg at launch, ELMO's innovative design cannot comply with the launch requirements, and it is therefore advised to perform further studies to re-evaluate the mission.

Strategic airlift is an essential capability for the rapid global movement of cargo, particularly in crisis operations where timelines are short and cargo demands extreme. Despite the European Union's collective purchase of A400M aircraft, capability gaps remain due to the aircraft's limited ability to airlift heavy and outsized equipment. This study applies Knowledge-Based Engineering aircraft design tools in conjunction with Agent-Based Simulation techniques to evaluate a fleet's ability to move cargo rapidly across strategic distances in high-stakes operational scenarios. One focus of the study is on the often-overlooked constraint of cargo hold volume and its effects on a fleet's ability to transport cargo. The study found that this volume constraint can reduce airlift capacity by roughly 20% in scenarios with medium-density cargo. Further, through design space exploration, this study identifies key top-level aircraft requirements for a new airlifter to work effectively in the European fleet of airlifters. Analysis reveals that a next-generation aircraft requires a wide fuselage capable of double-file loading, an extended fuselage length of 39m, an increased payload capacity of 120 tons, and a cruise speed of Mach 0.8. Such a design could enhance the fleet performance and reduce fleet requirements by roughly 30%.

Trajectory optimization has proven to be a powerful tool in solving a wide variety of optimal control problems in the aerospace field. However, in many cases, numerical complexities prevent the analysis of optimal trajectories for high-fidelity models, particularly due to the inherent difficulty of transcribing high-order dynamic systems. This research project proposes a methodology incorporating reduced-order modeling that retains the most critical dynamic characteristics from a full-order model while allowing the resulting simplification to be manageable for a trajectory optimization solver. The study applies this methodology to evaluate optimal landing trajectories for the UNIFIER19 C7A, a hybrid-electric aircraft equipped with a distributed electric propulsion system that was previously developed under the UNIFIER19 project. Results show that the reduced-order models generated for the aircraft can be used to generate flyable trajectories, verified by tracking the resulting landing approach paths using the base high-fidelity model. It is envisioned that this methodology will also be applicable to other aircraft models and mission phases.

As industries are trying to decrease their carbon footprint, new technologies are being implemented. This is also the case for the helicopter industry. Similarly to the automotive industry, the helicopter industry is looking to implement hybrid-electric propulsion. To support this, a lot of research has been conducted into the performance of the different hybrid-electric propulsion systems architectures.
However, no effort has been made to map the effects these architectures have on the safety of thehelicopter. Since safety is an important aspect of helicopter design, being able to grasp the effect ofvspecific architectural choices early in the development process could provide a major benefit in terms of development time. This knowledge gap led to the research goal addressed with this study: How is the risk analysis for helicopters affected by hybrid-electric engine architecture?

The safety assessment conducted during helicopter development is mandated via several industry standards, most notably SAE ARP 4754 and SAE ARP 4761. These standards were followed while performing a safety assessment for a baseline architecture provided by the thesis company. The first step was to identify the exact scope of the hybrid-electric propulsion system, explicitly determining which subsystems are considered part of the hybrid-electric propulsion system. The next step was the identification of the different functions that the system has to perform. By analyzing how these functions can fail several failure conditions could be specified. The failure conditions were then dissected until the failures could be traced to sub-system level failures. The knowledge gained from this analysis was then used to create a baseline to compare other theoretical architectures. The concepts were compared by failure rates for 5 different failure conditions, system weight, engine development requirements and
system complexity.

4 main classes of hybrid-electric propulsion system architectures were studied:
• Double-shaft Parallel
• Single-shaft Parallel
• Series-Parallel
• Series

Overall, it was found that choices in hybrid-electric propulsion system architecture significantly impact the safety assessment of helicopters. From the classes mentioned above, double-shaft parallel architecture has the least disadvantages, closely followed by single-shaft parallel. Series-parallel has higher failure rates but still shows a realistic possibility of implementation. Of all the architecture classes, the series architecture shows the worst results in all comparisons, lacking realistic implementation possibilities.

By combining this study with pre-existing performance studies and the recommended study on system weight, a comprehensive overview can be created to aid helicopter architects in the early developmentstages.

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.

One of the solutions for reducing emissions from the aviation industry is the implementation of electric or hybrid-electric propulsion systems in the aircraft. The German Aerospace Center (DLR) Institute of Electrified Aero Engines is investigating the feasibility of using fuel cells in hybrid-electric propulsion systems to power regional aircraft. The aim of the thesis is to develop a framework for the design, optimisation, and assessment of hydrogen fuel cell supply architectures for an Solid Oxide Fuel Cell (SOFC) powered aircraft. In addition to this, the implementation of waste reduction methods using a turbine is examined by modelling the air supply system. Key components for a hydrogen fuel cell supply architecture include cryogenic storage tanks, vaporisers, pipes, heat exchangers, pumps, turbines and compressors. Three architectures were defined using these components. Two of them utilised the tank self-pressurisation as the driving force behind the hydrogen flow in the system, with the one extracting gaseous and the other extracting the liquid hydrogen from the tank. The final architecture employed liquid hydrogen extraction using piston pumps. Analytical models for sizing each of these components and modelling the fluid flow through them were implemented in Python scripts. These were integrated in an optimisation loop utilising a local simplex-based algorithm within the RCE integration environment. Each architecture was optimised for two objective functions, the first being for system mass, and the second for both system mass and net power recovered. The optimisation parameters are the geometrical parameters of the components. Its constraints were related to dimensional limitations set by the aircraft and the fluid properties, such as pressure drops and fluid phases, throughout the system. Preliminary component sizings from previous project phases set the intialisation point. Results indicated that the choice of objective function greatly impacts the converged system. For the mass optimisation, the optimiser minimised themass of all the air side components. On the hydrogen side, a compromise is found between the mass of the component responsible for maintaining the pressure difference for the desired mass flowrate and the mass of the remaining components. When optimising for mass and power, the optimiser reduced air side pressure drops in order to maximise turbine power recovery and minimise compressor power, resulting in a heavier but more power-efficient system. The different architecture layouts primarily affected the system’s parasitic power, which includes additional power required to maintain tank and vaporiser storage pressure, as well as pump power. The implementation of a vaporiser in the liquid extraction architectures reduces this value, as less hydrogen needs to be vaporised in the tank and the mass of hydrogen in the vaporiser is less. Similarly, using a pump allowed for lower tank pressure, reducing both tank mass and power, leading to further reductions in parasitic power. These findings led to the selection of the pump-fed liquid hydrogen extraction architecture as the most suitable for the mission. Post-optimisation analyses included a sensitivity analysis to identify parameters with the largest impact on the converged system. The heat exchanger geometrical parameters had the largest impact due to their significant contribution to overall system mass and pressure losses, which affected other components like tanks and pipes. The analysis also shows that components related to the pressure drop, such as the pipes, become more important when optimising formass and power. A mission analysis for tank self-pressurisation in the gaseous extraction architecture revealed substantial additional energy required for tank operation, corresponding to a 14.7% system mass increase if stored in lithium-ion batteries. Finally, two additional sets of optimisations were performed with different initialisation points to examine its effect on the results. Although the converged systems differed from baseline optimisations, similar trends were observed.

Airline reserve fleet capacity is a strategic resource that provides airlines the means to add robustness to their schedules. This paper focuses on leveraging past data on airlines' execution of planned schedules (resulting delays, cancellations and missed connections), along with the disruptions that prohibited the flawless execution of the planned schedules, in order to assess the impact of reserve fleet composition changes on past days of operations. The fleet schedule is modeled as a set of parallel time-space networks, and a Mixed Integer Linear Programming model is defined and solved with the objective of minimising passengers' disruption costs incurred during execution of the planned schedule. By adjusting the reserve fleet within the model, the resulting impact on disruption costs and operational performance indicators is quantified. The approach presented can be applied on heterogeneous fleets and to hub-and-spoke airlines. Its performance is discussed using a case study on the network of a Full Service Carrier operating in Europe. Simulated results of all reserve fleet scenarios considered are compared with recorded passenger disruption data from the same period, demonstrating the model's capacity to emulate historically observed operational performance. Notably, the model (considering the same reserve fleet as was operated) exhibits an average deviation of 7.9% in disruption costs when compared with recorded data. The analysis of resulting disruption costs stemming from reserve fleet composition changes reveals that adding an A220 aircraft to the airline's reserve fleet would result in a slightly higher reduction (-5.7%) of disruption costs compared to adding an A320 aircraft (-5.3%). Conversely, removing an A320 aircraft from the reserve fleet would lead to a significantly higher increase in disruption costs (+20.5%) compared to removing one A220 aircraft as reserve (+16.9%). Consequently, when considering changes to the reserve fleet operated, the airline should prioritise alterations on the A220 portion of their fleet.

Hydrogen is seen as a potential energy carrier for the next generation of aircraft that will be more climate-friendly than the previous generation of kerosene-powered aircraft. Two types of hydrogen-based propulsion systems are currently foreseen for such aircraft: hydrogen combustion and hydrogen fuel cell. In addition to producing useful electrical power, fuel cell systems produce a considerable amount of heat which must be removed to ensure the continued and efficient operation of the fuel cell stack. This thesis presents a thermal management system sizing methodology for propulsive fuel cell systems onboard CS-23 commuter aircraft. The developed methodology is implemented into an aircraft sizing environment to study the characteristics of air-based thermal management systems and their effects on aircraft design and aircraft performance. The results show that the thermal management system, through its additional mass, parasitic drag and parasitic power, has both a direct and indirect effect on aircraft performance. In addition to the conventional thermal management system, an unconventional system using nanofluids is studied. The use of nanofluids showed no considerable improvement in both system and aircraft level performance when compared to the base fluid. This thesis shows the importance of considering the design of thermal management systems during the conceptual design of fuel cell aircraft.

This research project investigated the potential benefits of using hybrid hydrogen fuel-cell powertrains in commuter aircraft certified under the EASA CS-23 category to reduce the aviation industry's environmental impact. The study used an in-house aircraft design synthesis software to model conventional and fuel cell aircraft, including a sensitivity analysis. The fuel cell aircraft had a higher take-off mass, but almost equal mission energy use. Fuel cell aircraft brought a 76-98% reduction in global warming potential when using green hydrogen. The energy purchase costs were higher for fuel cell aircraft but equalized when emission allowance costs for the mid-term future were factored in. Overall, fuel cell powertrains have potential in reducing aviation's environmental impact, but further research is needed.

With the aviation sector growing each year, the need for a reduced climate impact is becoming increasingly important. Electrification of the propulsion system is believed to offer promising avenues in achieving this reduction. Additionally, airlines operating these aircraft have to adapt their operations and network to optimally utilize these aircraft. This research presents a methodology for the coupled design of a hybrid-electric aircraft fleet with strategic airline planning to optimally serve a specific network. The objective is to maximize the airline profit and minimize the network CO2 emissions. Aircraft design trade-offs in payload, range and runway length will guide the creation of new aircraft until the optimal aircraft fleet is determined. The methodology is tested in a case study for the regional airline network of SATA Air Acores. The study investigates the impact of introducing new hybrid-electric aircraft designs in the fleet on the creation of new aircraft, the aircraft allocation and the network performance. By directly integrating hybrid-electric aircraft design (having a parallel hybrid architecture) with strategic airline planning, it is possible to reduce the network CO2emissions by -11% at the cost of an airline profit decrease of -13%. When including a climate optimization, an additional reduction of network CO2 emissions is achieved of -27% with a small additional decrease in profit of -1%. Network profitability and climate impact are mainly dictated by the fleet diversity and the assumed technology level of the batteries employed in the aircraft. This research highlights the importance of including climate optimization in the design of new aircraft and the need for more advanced hybrid-electric propulsion architectures (such as distributed propulsion systems) to further contribute to climate impact reduction.

Interest grows rapidly in electric and hybrid electric aircraft. To determine the optimal performance and energy management required with such novel powertrain configurations, a knowledge-based aircraft and powertrain performance model is developed. The model is then used to set up an optimal control problem, which is transcribed to a non-linear programming problem using global orthogonal Legendre-Gauss-Radau collocation for single phase problems, and Hermite-Simpson local collocation for multiphase problems. The solution to the control problem allows identification of the best control strategies and energy management strategies. A case study is performed on the HY4 hybrid fuel cell aircraft and the hybrid electric Pipistrel Panthera. Solutions show that for best fuel economy, flying at a minimum drag airspeed, and keeping a constant power setting, proved more important than the choice of altitude. This was more noticeable for the HY4, with its relatively low power available and good aerodynamic properties following from its glider-based airframe.
The Fuel-optimal energy management strategies proved identical for both aircraft investigated. Batteries are used to provide a power boost during takeoff, after which batteries are discharged gradually throughout the remainder of the flight to maximize discharge efficiency. The engine or fuel cell are kept at approximately constant cruise power settings throughout the flight. The Panthera showed
consistent flight profiles with increasing range. For the HY4, however, achieved airspeeds reduced with increasing range, and additional measures were required to force a climb to non-zero altitudes due to its under-powered nature. The fuel-optimal trajectories offered an average of 10-15% of possible fuel
savings, depending mostly on the size of the onboard batteries. Fuel savings increased significantly at low ranges300km, where the contributions of the batteries have more impact.
Comparing different transcription methods and problem setups, it was concluded that global orthogonal, or pseudo-spectral, methods like Legendre-gauss-Radau collocation are not only faster, but also more consistent compared to simpler direct collocation methods. However, if the problem complexity increases and the performance limits of the aircraft are pushed, switching to a simpler method like Hermite-Simpson collocation reduced the time required to find a solution, with negligible differences in the resulting trajectories. Opting for a multiphase problem set-up, essentially splitting the problem in a series of individual subproblems, appeared less advantageous. While offering more control over the trajectories, time required to find solutions increased drastically, and offered no additional insight into the best energy management strategies.

The medium-altitude long-endurance (MALE) unmanned aerial vehicle (UAV) plays a key role in both military and civilian applications around the world. However, in a world where electric UAVs are increasingly becoming the norm, the MALE UAV is still primarily powered by hydrocarbon fuels and its design is very much constrained by existing general aviation reciprocating engines. The aim of this study is to investigate the design effects resulting from incorporating electric propulsion systems in MALE UAVs, while also investigating the effects of modelling fidelity on the results. Reciprocating engine, hydrogen fuel cell and battery models are incorporated into a multidisciplinary design optimisation framework, with the aim of minimising the maximum take-off weight at various levels of modelling fidelity. The results show that hydrogen fuel cells can provide around 40% reduction in the aircraft's maximum take-off weight, while batteries are not capable of enabling long-endurance flight. It is further shown that the effects of modelling fidelity of the aerodynamics and structural wing weight models are more significant when the aircraft design being optimised is substantially different from the baseline configuration.

The demand for air travel is increasing as more people gain access to commercial aviation. As the current generation of aircraft makes use of fossil fuel combustion, this growth results in an increase in the global emissions. This is at odds with the worldwide efforts of reducing the adverse effects of climate change. Therefore, advanced propulsion systems must be developed to limit the emissions caused by the commercial aerospace industry. Using hydrogen fuel cells for propulsion is a promising technology to potentially get the sector to zero emissions. It is of interest to explore the capabilities and feasibility of aircraft with a hydrogen fuel cell powertrain. To determine the feasibility for a wide range of aircraft, a general design methodology is required.Current research efforts focus on component level performance, however system level design research, while present, didn't introduce a general methodology. The most pressing challenge was found to be related to the system level design of a CS-23 category hydrogen fuel cell aircraft. The CS-23 category aircraft class has been identified as the most suitable focus for research efforts, due to the lower technical and certification requirements placed on the components to reach a feasible design. A general methodology for the design of CS-23 category aircraft was therefore found to be a useful contribution to the state of the art. This additionally provides a deeper understanding into the most important parameters of the power and propulsion systems design.In this report, a general methodology for the conceptual design of hydrogen fuel cell powered CS-23 category aircraft is presented. The methodology makes use of a modified class 1 weight estimation for initial sizing according to customer requirements and component technology levels. The generated aircraft is refined using further aerodynamic analysis, which results in a feasible aircraft concept, as well as component level specifications for important aircraft components.The methodology is implemented in a software tool, HAPPIE (Hydrogen Aircraft Power \& Propulsion Initial Estimator), which allows for rapid sizing of different hydrogen fuel cell concepts. This makes the methodology accessible, and additionally provides feedback on the effects of individual design choices and technology levels on system level performance. SUAVE is used to perform the refined aerodynamic analysis.The methodology is validated using existing conventionally powered aircraft, by comparing the sizing results from the methodology with publicly available data. The methodology's ability to analyse conventional as well as hydrogen fuel cell powertrains, furthermore allows for performance comparison between current and future technologies.The results of the sizing methodology demonstrate the viability of hydrogen fuel cell aircraft in the CS-23 category. A conceptual design is generated, which serves as a baseline for the sensitivity analyses. It is found that hydrogen fuel cell aircraft are generally heavier than conventional aircraft, using current technology levels. Liquid hydrogen is identified as the best hydrogen storage method. Compressed hydrogen storage is also possible, however this results in a heavier aircraft with limited range. The current methodology does not predict thermal behaviour to have a significant effect on the mass of the aircraft. A component sensitivity analysis determined that the fuel cell efficiency, fuel cell specific power and hydrogen storage efficiency are the most important parameters.The ideal cruising altitude for fuel cell aircraft is at an intermediate altitude, due to the fact that the fuel cell powertrain performance decreases at increasing altitude, which balances with the lower drag in lower density air.The research demonstrates that hydrogen powered CS-23 aircraft are viable for current technology levels, and are a suitable way in reducing carbon emissions in this category.

Classically, aircraft controls are designed such that every control surface type primarily influences a single degree of freedom by creating a moment. Increased availability of computational resources and novel aircraft configuration allow a deviation from this approach and to utilize individual control surfaces to generate moments around multiple axes. The research investigates the impact of control allocation algorithms on the required control surface span and area for a box-wing configuration aircraft, the PrandtlPlane. The unique geometry of two full wings allows more flexibility in control surface placement. An optimization system for automatic control surface sizing under the constraints of adequate handling qualities has been developed and used to compare mechanical gearing, the constrained pseudo inverse, and the direct allocation algorithms. The results show that the PrandtlPlane configuration can benefit from the use of control allocation, showing a clear advantage of the direct allocation algorithm.

Fuel cells are one of the promising technical solutions for the first generation of sustainable aircraft. Currently, there are several shortcomings in the literature related to the conceptual and preliminary sizing of the propulsion systems of such aircraft: not only is there a lack of general methods, but also, much less effort is put into the estimation of the system’s weight, as compared to its performance. In this thesis project, a set of preliminary sizing methods was developed for the propulsion systems of CS-25 proton-exchange membrane (PEM) FC aircraft. These methods were then integrated into the Initiator, an aircraft sizing environment. With this approach, a complete preliminary sizing procedure of a regional turboprop aircraft was successfully performed. Finally, sensitivity studies were carried out in order to demonstrate the advantages of this modelling approach.

This research is about the integration of hybrid and electric aircraft (HEA) in the air
traffic management system. Current conventional aircraft are responsible for emissions and noise that often lead to nuisance for residents. In order to make aviation more sustainable, aircraft manufacturers are studying the possibilities to replace fuel with a battery. However, due to a different powertrain, the performances deviate from conventional aircraft. This affects the air traffic management operations. This research studies how to integrate hybrid and electric aircraft in between conventional aircraft at (crowded) airports.

It is the aim of this research to assess the mission performance of a boxwing aircraft by developing a configurationagnostic, multifidelity optimal control toolbox for performance and mission analysis. The boxwing aircraft, sometimes named a PrandtlPlane (PrP), is an unconventional aircraft. An instance with redundant controls is designed within the PARSIFAL project. This specific aircraft is designed for commercial transport in the shortrange segment (a 4000 km design range) and for a high passenger capacity (up to 308 passengers). Because of the beneficial induced drag characteristics inherent to the boxwing configuration, the PrP represents a possible solution towards the sustainable future of aviation. To investigate the potential of the PrP as an alternative to conventional commercial aircraft, the mission performance assessment of the aircraft has been split into two components. The first part of the assessment covers the comparison between the performance of the PARSIFALdesigned PrP and that of a competitor aircraft with a similar design range, the A320, while allowing only nonredundant controls. The second part of the assessment involves the quantification of the PrP’s performance when allowing redundant controls in the form of Direct Lift Control (DLC), enabling the aircraft to increase its net lift without a change in pitching moment. Analyses of the PrP and its competitor aircraft for various ranges have shown that the PrP outperforms its competitor in terms of relative fuel consumption. When flying its minimumfuel mission, the PrP’s competitor consumes less fuel in absolute terms. Nonetheless, because the PrP carries more than twice as many passengers, it consumes up to 14.5 % less fuel per passenger per kilometre. In other respects the PrP’s performance is inferior to that of its competitor. The 5400 km maximum range of the PrP is considerably lower than its competitor’s maximum range of 6200 km. Moreover, at a fueloptimal Mach number of approximately 0.7 the PrP cruises appreciably slower than the cruise Mach number for which it was designed, unlike its competitor. In general, the PrP flies its trajectories much slower than its competitor at an approximately 10 % lower average velocity in the minimumfuel missions. If both time and fuel are considered equally in the cruise altitude optimisation, the design altitude of 11 km is deemed appropriate. If only fuel consumption is considered, the PrP would benefit in fuel economy from lowering the initial cruise altitude at the cost of increased mission time. At an optimal altitude of 9.3 km, the PrP would consume 2.2 % less fuel than at its design altitude of 11 km at the cost of even slower flight. The sensitivities of the PrP’s mission time and fuel performance to changes in its design Zerofuel Mass (ZFM) have been investigated. Keeping the Maximum Takeoff Mass (MTOM) constant while varying the ZFM, design mission simulations were run for the PrP for several objective functions. It was found that when flying for minimum fuel, a 1 % increase in ZFM incurs a fuel consumption penalty of over 1 % through a nearlinear, direct proportionality. Likewise, the mission time varies nearly linearly with the ZFM; a 1 % increase results in an approximate mission time increase of nearly 0.5 %. The incremental aerodynamic lift and drag due to control surface deflections for DLC were modelled using a flatplate approximation. With this approach, the projected missionlevel benefits of using DLC are marginal. On the design mission, the results indicate an increase in fuel economy of 0.6 % on the minimumfuel mission and negligible temporal gains on the minimumtime mission. It is however emphasised that numerical uncertainties due to the discretisation of the problem pollute all obtained solutions to some degree, such that appropriate caution should be exercised when interpreting these results in an absolute sense. In future research, a grid refinement study would be a valuable addition to quantify and bound these uncertainties. It is deemed equally important to look into a more sophisticated way to model the control surface aerodynamics necessary for assessing the benefits of DLC. A broader recommendation pertains to future research on boxwing aircraft aerodynamic design. The current research has indicated that the optimal trajectories for the PrP result in very distinct flight profiles when optimising for different objectives. Therefore, it would be interesting to see how the aerodynamic design could evolve, such that flying for fuel economy wouldn’t require such a compromise in temporal performance and vice versa.

With the global increase in passenger traffic and growing popularity of long-haul routes over the Asia Pacific region and Atlantic Ocean, the possibility for hypersonic transport could become an attractive option to reduce flight time over long distance from 16-20 hours down to around 4-5 hours. In this thesis, a Multi-Disciplinary Optimisation platform has been developed to allow for the optimal sizing of hypersonic transport vehicles using vehicle take-off mass as the performance indicator subjected to fuel volume and payload height constrains. The current platform is applied to the LAPCAT A2 hypersonic long-range transport configuration by Reaction Engines, to determine the impact of range and cruise Mach number on the design of hypersonic aircraft. Results show that the optimal shape is greatly dependent on the aircraft range and fuel volume constraint. Additionally, the optimum hypersonic cruise Mach number is dictated by a trade-off between mission time, engine efficiency and Thermal Protection System mass.