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.

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.

The aviation sector and its associated activities contribute to climate change, damage to the environment, community noise, and local air pollution. As a consequence it affects the health and quality of life of citizens that live close to airports, and through climate change every person on earth. To reduce the impact of air travel and thereby accelerate growth, new technologies are being investigated. One area of research looks into the opportunities of using batteries and electric motors as additional energy source and power converter. Besides solving some of the current problems, these new technologies could enable radical new forms of aviation and create new business models such as on-demand personalized air travel. Because the design rules of novel hybrid-electric aircraft differ significantly compared to that of conventional aircraft, new methods must be developed tailored for this new technology. Due to the additional variables introduced by using hybrid-electric propulsion systems, a simple sizing study quickly becomes a multi-disciplinary optimization problem. Furthermore, these new propulsion systems should not be benchmarked in an isolated way due to synergistic benefits with other disciplines. To close the gap between clean sheet designs and retrofits, it has been concluded that research focusing on feasible concepts in the near future is required. The objective of this research is to define the implications of using hybrid electric propulsion systems in general aviation aircraft. The influence of design, operational and mission choices on optimal performance are identified through optimization studies. The design space includes the geometry of the main wing and the operational variables of the propulsion system. The performance is either measured in terms of fuel or total energy consumption per kilometer for design ranges between 50 and 1000 km. To perform the quantitative analyses, a conceptual design framework has been developed. Models of all propulsion systems have been developed that simulate their performance. Every part of the tool is either verified with experimental data or has already been verified in literature. A quasi-three-dimensional aerodynamic solver has been modified to determine the aerodynamic characteristics with minimal computational time. The battery performance is modeled as function of time and rapidly sized such that the required power can be delivered by the smallest battery possible. A multi-disciplinary optimization approach has been used to integrate all modules and converge to an optimal design as quickly as possible. First of all, a difference is found in terms of optimal configuration and operation of an aircraft when designed specifically for minimal fuel or for minimal total energy, leading to a maximum difference of 5% in terms of energy consumption. A trade-off between designing for aerodynamic efficiency versus the ability to carry batteries is part of the optimization routine. It has been shown that aircraft optimized for fuel consumption carry systematically more batteries compared to aircraft optimized for total energy consumption, always leading to a heavier aircraft with larger wings. Regardless of objective, the aerodynamic efficiency becomes more important for an increasing range while the ability to carry batteries decreases. Three variables have been defined that determine the amount of hybridization in terms of power and energy without the need for constraint functions, as any combination of these variables lead an inherently feasible design: a climb coefficient that determines the rate of climb as a fraction of the maximum rate of climb, the fraction of the cruise that needs to be performed full electric, and a hybridization factor that determines the throttle of the internal combustion engine. Given a certain required power at the power management system, there exist one specific throttle that leads to the maximum effective efficiency of the internal combustion engine. This throttle depends on the round trip efficiency of the battery charge-discharge cycle. When the internal combustion engine is scalable, the optimizer always makes sure that the most efficient throttle matches the power requirements during cruise, i.e. it avoids recharging. The maximum possible range is highly determined by the required cruise velocity, given a fixed internal combustion engine. By scaling the internal combustion engine such that it is powerful enough to deliver the required power for cruise, the maximum range can be significantly increased. Furthermore, the optimal cruise altitude is found to be increasing with range and cruise velocity. At a cruise speed of 60 m/s the optimal cruise altitude coincides with the lower bound of 2 km whereas the optimal altitude at a cruise speed of 90 m/s increases from 2 km to 4 km for design ranges from 350 km to 700 km. In the latter case, the cruise altitude settles at 4 km for longer ranges. The maximum range at which full electric cruise is possible is found to be the most efficient range in terms of energy consumption. This range increases linearly with the battery specific energy, providing designers a good initial estimation tool. This research adds to the scientific body of knowledge as it presents a method that solves the multi-disciplinary optimization problem associated with the design of hybrid electric aircraft. Furthermore, through multiple optimization studies it provides insight in the influence of the design choices, operational choices, constraints, and mission profiles on the optimal performance of general aviation hybrid electric aircraft.

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.

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.

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.

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.

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.

Considerations that decide producibility of a design are an important part of the design process, and must be included in the early design stages to ensure that these designs can be realised. Not including these considerations carries the risk of incurring additional costs and delays at later stages of product development because of design changes, or can lead to limiting oneself to conservative design choices to reduce the associated risk. The current process in the industry accounts for these production considerations in design through a manual process that is iterative and time-consuming, and hence forms a bottleneck in being able to trade-off multiple design concepts. Attempts at accounting for these production considerations in an automated way are associated with the limitations of either only considering the manufacturing cost, being specific solutions that work only in certain scenarios, or being dependent on some commercial software tools, which are not fully suitable for use in context of automation and/or at the conceptual design stage. Additionally, the aspects of manufacturing and assembly are usually not considered at the same time in these studies. Therefore, this thesis aims at developing a methodology that enables the automated inclusion of production considerations in the conceptual design process of aircraft structures, while overcoming shortcomings of the state-of-the-art....

The Flying-V, a novel aircraft design developed at Delft University of Technology, presents a revolutionary flying wing transport aircraft with a remarkable 20% reduction in energy consumption compared to traditional twin-aisle planes. This thesis project delves into the study of optimizing the Flying-V's landing performance, emphasizing the necessity of reducing pitch attitude. High-lift devices, particularly split flaps, were explored for this purpose. Wind tunnel tests were carried out on a scaled-down model of the half- wing, in the Open Jet Facility of TU Delft. The tests yielded two successful flap configurations— a single-flap and a double-flap. These were analyzed further using a flight performance tool to make a final selection on the flap configuration. The single-flap option proved effective in reducing landing pitch attitude by 3 degrees, significantly lowering obscured segment by 20 to 30 m and the pilot's eye altitude by 1 m. This is a quite desirable outcome for the landing performance of the Flying-V which significantly improves pilot’s vision.

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.

Developments in computational capabilities and the always increasing demand for higher performance of internal flow applications has meant that Computational Fluid Dynamics (CFD) has become an essential tool within the design process. A key point of interest is to couple the fluid solver with numerical optimisation techniques in order to obtain a more automated design process that can handle a vast amount of different design aspects. The open source CFD suite SU2 has emerged as an enabler for aerodynamic shape optimisation involving a large number of design variables, due to its efficient, accurate and flexible discrete adjoint solver.  For the adjoint-based aerodynamic design optimisation of internal flow applications the deformation of the volumetric mesh has to be performed in an robust and efficient manner. Often small wall clearance gaps and periodic domains are encountered in internal flow domains, which could potentially lead to the deterioration of the mesh. Sliding boundary node methods can be applied in order to maintain the mesh quality in case of small wall clearance gaps. Additionally, periodic boundaries can be displaced in a periodic manner following the applied deformation in order to prevent low quality cells near the periodic interface. Therefore, it would be of interest to implement the sliding boundary node methods and periodic conditions in a Radial Basis Function (RBF) interpolation method, one of the most robust mesh deformation methods available.  Additionally, the computational efficiency should be considered, since high computational times should be prevented for large and complex three-dimensional cases with a high number of design variables.  The aim of this thesis project is therefore to develop a robust and computationally efficient mesh deformation method suitable within the discrete adjoint optimisation framework of SU2 for internal flow applications by means of developing an implementation of the RBF interpolation method including sliding boundary node algorithms, periodic boundary conditions and data reductions methods. The sliding is achieved by replacing the interpolation condition for the sliding nodes with a planar slip condition. Or alternatively, by freely displacing the sliding nodes based on the known deformation and subsequently projecting the nodes back onto the boundary. The periodic displacement of the boundaries is ensured by making the distance function of the RBF periodic. The periodic nodes are then treated as internal nodes to allow them to move. The developed RBF-SliDe tool is able to generate higher minimum mesh qualities compared with the regular RBF interpolation method. The sliding of the boundary nodes reduces the degree of skewing of the mesh elements in case of drastic deformations, resulting in a higher minimum mesh quality. Furthermore, the introduction of the periodic displacement prevents low quality skewed or compressed mesh elements, as the periodic boundaries move along with the deformation.  The Aachen turbine stator blade is considered as a realistic three-dimensional test case. For this stator blade an optimised geometry was available, which was obtained with an adjoint-based aerodynamic optimisation performed with SU2. Therefore, the resulting minimum mesh quality is compared to the one obtained with the more conventional linear elasticity equation method as used in SU2. The minimum mesh quality obtained with the RBF-SliDe tool is nearly three times higher compared to the minimum mesh quality of the linear elasticity equations methods. This highlights the potential of the periodic sliding RBF interpolation method in terms of preserving the mesh quality.

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.

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.

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.

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.

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.