Hybrid aircraft are a necessary step in the energy transition towards sustainable aviation, given current limitations in battery technology. However, hybrid systems face significant economic and technological hurdles of their own. This study investigates whether designing hybrid-electric aircraft as part of a family—with shared components and design commonality—can offset economic drawbacks while maintaining environmental benefits. A novel methodology was developed and validated against existing literature, integrating economic and climate models within a hybrid aircraft family design framework. This approach facilitates aircraft family design by strategically constraining design freedom at the subsystem level, introduces a refined calculation of commonality indices, and incorporates these indices into a bespoke economic evaluation framework specifically tailored for commercial hybrid aircraft. Results from a parametric case study demonstrate that increased commonality yields drastic improvements in economic feasibility —on the order of billions of dollars—, while environmental performance varies by less than5%across design families with differing levels of commonality. These findings underscore the critical role of commonality in determining program viability, an effect not previously quantified for hybrid systems. Additionally, family design trends influence program value and emissions in nuanced ways. This work provides guidance for optimizing hybrid aircraft family designs and highlights opportunities for further research in sustainable and economically viable aviation.

Due to climate change and rising sea levels, a sustainable formof airborne cargo transport is needed for quick disaster response in rough sea conditions.

The increasing urgency to mitigate climate change has underscored the need to transition from conventional fossil-based aviation fuels, such as kerosene, to sustainable alternatives. Hydrogen stands out due to its potential to significantly reduce greenhouse gas emissions, making it a promising energy carrier for the aviation sector. However, adopting hydrogen presents substantial challenges, with the development of specialized fuel containment systems being one of the foremost obstacles. Double-walled tanks employing vacuum insulation offer an effective solution for cryogenic hydrogen storage, but they require a robust supporting structure for the inner vessel. Various solutions have been proposed to support the inner vessel of cryogenic tanks; however, despite the diversity of designs, there remains a noticeable lack of comprehensive research focusing specifically on the structural behaviour and feasibility of these inner vessel supporting structures for commercial aircraft applications. This study proposes a novel fibre-based suspension technique for the inner vessel of a double-walled integral tank designed for liquid hydrogen storage in large commercial aircraft. A finite element model was developed to evaluate the structural interaction between the inner and outer vessels and the supporting fibres, enabling structural sizing optimization to assess the impact of added loads. A parametric study was conducted to explore the influence of fibre design parameters on structural performance, mass, displacement, and thermal behaviour. Key design guidelines were established. First, using more than two longitudinal anchoring points results in the unwanted transfer of bending loads from the outer to the inner vessel. Second, while increasing the number of circumferential fibres reduces peak loads and displacements, the associated anchoring mass is the primary limiting factor, as thermal conduction was found to be negligible. Lastly, fibre orientation should prioritize low stiffness in the contraction direction to minimize tensile forces under initial filling. This should be combined with fibres angled in the longitudinal direction to improve longitudinal stiffness and displacement control. The results confirm the structural feasibility of the suspension system, showing only a marginal structural mass increase of approximately 1.88 % compared to a baseline integral tank without internal support. These findings provide practical guidance for the implementation of fibre-based suspension systems in cryogenic tank structures, supporting the development of hydrogen storage solutions for aviation.

The Delft Laminar Hump (DeLaH), discovered by the Aerodynamics Department of the Faculty of Aerospace Engineering at the Delft University of Technology, is a symmetrical smooth hump that is placed at a set distance parallel to the leading edge of the wing that reduces skin friction drag as it attenuates the growth of the crossflow instabilities (CFI). With the hump, transition can be delayed up to 14%. The only requirements are that the hump needs an natural laminar flow (NLF) airfoil and a condition where CFI is the dominant transition mechanism to be effective. Generally, CFI dominates when the wing sweep is greater than 30 − 35 deg, making the vertical stabilizer the best candidate for implementing the hump.

This research implements the Delft Laminar Hump on the vertical stabilizer of subsonic transport aircraft by modeling the effect of the hump as a shift in transition location. By using a Quasi-3D aerodynamic analysis in combination with a transition location database, the effect of the hump on the lift and drag coefficient of the vertical stabilizer is analyzed. The transition location database is constructed by using the external velocity of airfoil sections of the vertical tailplane and the boundary layer solver and stability analysis developed by the Group of Flow Control and Stability within the Delft University of Technology. With this, the N-factor curves along the chord can be calculated. Knowing the respective N-factors of the clean and hump configuration, the associated transition locations can be determined, which are then used to calculate the lift and drag coefficients of the vertical tailplane. To evaluate the aerodynamic effect of the hump on the full aircraft directional and lateral stability a stability analysis based on a method by Fokker / Obert is performed and checked against the CS-25 for Large Aeroplanes regulations by European Aviation Safety Authority (EASA).

It was found that the Delft Laminar Hump (DeLaH) on the vertical stabilizer of a subsonic transport aircraft does not affect the vertical tailplane lift curve slope, thus not affecting the stability of the aircraft. In contrast, the vertical tailplane drag coefficient is reduced by the hump. Retrofitting the hump on Airbus A320 (conventional tail) and the Fokker F-28 Mk1000 (T-tail) results in a reduction of the vertical tailplane drag coefficient of 6.73% and 8.72%, respectively. Translating this vertical tail drag reduction to the full aircraft drag coefficient results in a reduction of 0.17% and 0.34%. To evaluate the effect of the hump on weight and fuel consumption, additional weight and mission analyses are performed. Evaluating the harmonic range, an fuel reduction due to the hump of 0.16% and 0.32% is established, for the Airbus A320 and Fokker F-28 Mk1000, respectively. The aircraft weight is reduced by the same percentage through the fuel reduction, as it is assumed that the added weight due to the hump itself is negligible.

Additionally, two sensitivity analyses were performed, namely, sweep angle variation and surface area scaling to analyze the effectiveness of the hump for different vertical tailplane geometries. The effect of sweep angle on hump effectiveness does not affect the vertical tailplane lift curve slope and thus also not the stability coefficients. For the full aircraft drag coefficient, the hump effectiveness has an exponential relation with sweep angle and is most effective at lower sweep angles. A maximum full aircraft drag reduction was found of 0.41% at 30 deg sweep with an equivalent fuel reduction of 0.39%. Overall it is concluded that lower sweep angles are beneficial as the hump is most effective and results in reduced vertical tailplane weight, less fuel weight as well as an increased stability margin. Analyzing the effect of surface area scaling, the hump has no effect on the vertical tailplane lift curve slope regardless of surface area, again retaining the aircraft’s stability. For the full aircraft drag coefficient, the hump effectiveness increases linearly for increasing surface area up to 0.37% at a surface scaling factor of 1.2 times the original vertical tailplane surface area. In terms of fuel reduction, a maximum value of 0.35% was found. There will be an optimal vertical tailplane surface area, since the hump effectiveness increases for increasing surface area, whilst for the full aircraft drag, vertical tailplane weight, and fuel weight a smaller surface area is preferred. The stability margin becomes the limiting factor as a minimum surface area is required for sufficient stability. Comparing the baseline aircraft, the hump is more effective for the Fokker F-28 Mk1000 over the entire range of scaling factors and sweep angles. This leads to the suspicion that taper- and aspect ratio, and cruise speed play an important role in the effectiveness of the hump, but more research is required.

This research shows that the Delft Laminar Hump has a significant drag-reducing effect on the overall aircraft, whilst not affecting the aircraft’s stability. Even though the fuel savings for an individual aircraft are not very large, on a fleet level this would be significant. The hump can be retrofitted on existing aircraft by gluing it on the outer skin, making it a relatively simple and cheap way to improve efficiency for aircraft manufacturers. Nevertheless, more research is required before implementing the hump on commercial aircraft as it is still unknown whether the hump also works on the suction side of the wing as well as whether the hump causes a shock at cruise Mach. Also, interaction effects with the horizontal stabilizer and fuselage need to be taken into account, to fully quantify the effectiveness of the hump. Dedicated wind tunnel experiments, flight tests, and/or CFD simulations are necessary.

The aviation industry is currently facing significant pressure to enhance its sustainability by increasing aircraft energy efficiency and reducing its climate impact. A promising approach to fulfilling these demands is to improve the aircraft's aerodynamic performance through drag reduction by implementing laminar flow technologies, particularly Natural Laminar Flow (NLF) or Hybrid Laminar Flow Control (HLFC).
Prior works assessing laminar flow technologies have mostly focused on evaluating their aerodynamic performance and, in the case of the HLFC, on the influence of system design. The impact of these technologies on the overall aircraft performance has received only limited consideration, with the majority of studies focusing on long-range aircraft, utilizing simplified models for HLFC systems, and considering only one laminar flow technology at a time.
This study adopts a holistic approach to assess the potential fuel savings that could be achieved by combined application of NLF and HLFC technologies on the various components of a short-to-medium range aircraft concept, with an intended entry into service in 2035. To achieve this objective, a conceptual aircraft design process is employed. This process captures the aerodynamic effects of laminar flow technologies and fully integrates the HLFC system design to provide an accurate estimate of aircraft performance.
The findings of this study reveal a potential for fuel savings of 5.9\% on the design mission through the combined application of NLF and HLFC, compared to a turbulent aircraft with an equivalent technology level. Additionally, the results indicate that strategic combination of the two technologies on a single component can significantly reduce complexity while further enhancing fuel savings. A failure analysis also provides an initial estimate of the impact of various failure scenarios on the aircraft's performance.
These findings demonstrate that, despite the aircraft's short range, the combined implementation of the two laminar flow technologies offers a potential for fuel savings with reduced complexity, motivating further research in their application to this aircraft category.

Within the context of the new wave of eVTOL aircraft development, Cyclotech GmbH has been working on the development of cyclorotors, a type of propeller consisting of rotating blades along a common longitudinal axis, where the pitch angle of the blades is actively controlled through what is known as a pitch curve. This thesis proposes a methodology to solve an inverse design of this pitch curve, which results in a certain target lift distribution. A framework using Bayesian inference is proposed, altogether with CFD simulations for the aerodynamic solution of the cyclorotor. To reduce the number of CFD evaluations and improve the framework efficiency, gradient descent methods are used altogether with semi-transient adjoint CFD gradients. Furthermore, low-fidelity information is added to the framework through the definition of a prior model to help speed up the inference. Using this methodology, three different test cases, from lower to higher complexity, are defined and successfully solved.

Growing concerns about the environmental impact of aviation have sparked interest in hydrogen aircraft as a greener alternative. Hydrogen can be used to power existing turbofan engines or electrical motors via a fuel cell, eliminating carbon emissions not only during flight, but also during production, provided renewable energy sources are used. However, adopting hydrogen as fuel introduces technological challenges, particularly with regard to on-board storage. Integral tanks, which are part of the aircraft's main structure, seem promising but existing designs show limitations in their integration with the airframe and insulation capabilities.

To address these issues, this study proposes an integral tank concept featuring a double wall architecture with vacuum insulation. The main advantage of this design is the use of an external stiffened wall that can be directly connected to the remaining airframe. In addition, having stiffeners on the outside ensures the required space for systems routing and addresses concerns with the crash worthiness of the structure. A parametric method, coupled with finite element analysis is developed to size the external load bearing wall, enabling quick analysis and mass estimations of different tank configurations. The method consists of a sizing optimization with the objective of minimizing the structural mass under constraints on the strength, buckling stability and fatigue behaviour.

The feasibility of the concept is then evaluated on an aft tank for a short/medium range aircraft in configurations with and without a forward tank. Preliminary results under this realistic scenario point to fuel containment efficiencies of up to 0.71, which are consistent with existing designs. Moreover, buckling stability is identified as the critical design criterion, highlighting the importance of using a stiffened shell design. These findings show the viability of the proposed concept from a structural standpoint and provide the basis for further research. The optimum solution at an aircraft level can be obtained by integrating the developed framework into a multidisciplinary aircraft design tool.

With the drive to decarbonise the motorsport industry a variety of clean powertrains were developped and raced. While extensive lap time optimisation has been performed on hybrid and battery electric cars, no such optimisation has been performed on hydrogen fuel cell racing cars yet. This work models the Forze 8, a hydrogen fuel cell electric racing car, and compares its optimal lap, found by solving optimal control problems, to racing cars equipped with battery electric and internal combustion engine powertrains, all with the same peak power and weight on two different tracks. In qualifying scenarios, the battery electric racing car is faster than the internal combustion engine car by 1.0 [s] on both tracks, while the fuel cell electric car is slowest, trailing 5.3 and 3.8 [s] behind the battery electric car at Assen and Zandvoort respectively. In racing scenarios, the internal combustion engine car is quickest with the fuel cell electric car behind by 5.6 and 2.9 [s] at Assen and Zandvoort respectively. The battery electric car has the best single-lap performance due to its high continuous power, while the fuel-cell electric car is better suited for endurance scenarios as it carries more energy. This work shows that both battery and fuel cell electric racing cars can be valid alternatives for internal combustion engine racing cars.

Due to the high aspect ratio and low induced drag of aircraft with strut-braced wings, they are being extensively studied due to their fuel-saving potential. This thesis aims to expand the fundamental knowledge in the design of aircraft with Strut Braced Wings (SBW) by achieving the following objectives.

The objectives of this thesis are twofold, with their primary focus on evaluating the effect of changing the typical design variables for an SBW, such as the wing span, root chord, spanwise strut attachment location, wing taper ratio, wing sweep and engine location on the aerodynamics, weights and performance of an SBW. The first objective was to evaluate the significance of including the propeller slipstream effects in the preliminary stage design optimisation of a low-speed, short-range SBW. This aligned with the hypothesis that the performance of an SBW could be enhanced by using swirl recovery. The second objective of this thesis was to investigate the sensitivity of various design variables to the aircraft’s fuel burn and other performance metrics at the optimum SBW design.

A Design of Experiments (DOE) approach was used to explore the design space involving the typical influential SBW design parameters. The geometry and mesh files were created using OpenVSP for every DOE point. This was followed by the aerodynamic analysis in a panel method-based software called Flightstream, which could capture the relevant aerodynamic flow phenomena with reasonable accuracy. The aerodynamic analysis was performed twice- first without considering slipstream effects and second by simulating them. Regression-based analytical equations, particularly designed for the weight estimation of the wing and strut, were used from the literature. Empirical equations from FLOPS were used for the rest of the aircraft components. The performance of the SBW was calculated iteratively using the Breguet range equation along with a few modifications. All the SBW designs were constrained by a maximum wing loading criterion.

From two sets of the DOE results (with and without propeller effects), it was observed that the two optimum SBW designs were identical in terms of external geometry and performance. A deeper investigation revealed that the variation in the spanwise engine positioning (to maximise the swirl recovery) resulted in a marginal change of the induced drag (less than two drag counts). It was concluded that the propeller slipstream effects could be excluded from the preliminary stage, design optimisation of a propeller-powered, short-range SBW to reduce computational expense. However, the results should be treated with a pinch of salt due to the limitations of panel methods. Moreover, the SBW was optimised only for cruise, not for other flight phases such as take-off and climb, which may benefit from swirl recovery.

Finally, a sensitivity analysis was performed to evaluate the trends in the performance metrics, such as fuel burn, lift-to-drag ratio, wing loading and maximum take-off weight when subjected to variations in the design variables at the optimum. The impact of the fuel burn was quantified, while the other performance metrics were qualitatively answered. The research findings revealed that certain design variables had a greater influence on fuel burn than others when varied by ±10% w.r.t their optimum value...

The potential for a correction by Rethorst for Vortex Lattice Methods (VLMs) exists to improve Longitudinal Static Stability predictions when applying the Rethorst correction to a slipstream engulfed horizontal stabilizer. The Rethorst correction has already proven to result in accurate lift predictions in VLMs for an isolated wing in a slipstream jet. The aim of this thesis is to investigate the applicability for the Rethorst correction at the horizontal stabilizer, by evaluating the derivation assumption and attempting to obtain an accurate prediction of the stabilizer lift using the correction. With this regard, this thesis has been unsuccessful. An overestimation in the average dynamic pressure and downwash are obtained at the stabilizer, which possibly result from modeling issues. The left over work necessary to successfully draw a conclusion is identified and involves investigating the possible modeling issues and the Rethorst correction application outside its assumption scope.

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.

Boundary Layer Ingestion (BLI) is a promising propulsion concept for aviation, aiming to reduce fuel burn. This configuration involves using the propulsor to ingest the boundary layer from the fuselage or wing. Typically, aircraft with BLI propulsors also integrate a hybrid-electric powertrain. This integrated setup has shown potential in fuel burn reduction. This thesis investigates the interactive effects of BLI-induced power-saving benefits and aircraft design parameters on overall performance, focusing on a fuselage tail-mounted BLI propeller.

The thesis has three primary objectives. First, it seeks to enhance the fidelity of the existing BLI model within a conceptual aircraft design tool. The improvement involves transitioning from the actuator disk theory to the blade element theory (BET) with gradient-based optimization. Additionally, a surrogate model is established with this improvement through multidisciplinary design optimization (MDO), design of experiment (DoE), and response surface methodology (RSM) to predict power-saving benefits based on fuselage geometric and operational parameters. Comparison reveals discrepancies between the existing BLI model and the surrogate model, emphasizing the influence of blade aerodynamics.

The second and third objectives delve into the sensitivity of aircraft-level performance and powertrain settings. The conceptual aircraft design tool, Aircraft Design Initiator (Initiator), is used for sizing, with the surrogate model integrated. The study uses a regional turboprop ATR-72 as a reference conventional aircraft design and employs a partial-turboelectric (PTE) architecture for the hybrid-electric powertrain in the radical aircraft design. A sensitivity analysis varying design parameters by 20%, including fuselage slenderness ratio, propeller size ratio, and shaft-power ratio, reveals their impact on BLI effects and aircraft-level performance.

The study reveals that fuselage length significantly impacts fuel weight and, consequently, aircraft performance. Surprisingly, aero-propulsive benefits from BLI do not directly enhance overall aircraft performance, attributed to an associated weight penalty. Instead, the shaft-power ratio and propeller size ratio prove significant at the system level. Further investigation into powertrain settings suggests that radical BLI-equipped aircraft designs are not able to surpass the energy efficiency of conventional counterparts. Additionally, the proposed surrogate model exhibits a limited applicable range, especially under high BLI propeller disk loading, rooted in underlying physical constraints in propeller design. These findings prompt a critical examination of the feasibility of hybrid-electric powertrains with BLI for regional turboprop aircraft, with a note of caution regarding potential variations in modeling methods and assumptions.

Aircraft design systems are frequently used to synthesize aircraft designs. However, it is inherently difficult to reconfigure these software systems to facilitate a broader range of design studies (e.g. optimization or sensitivity studies) and to address follow-up questions about the synthesized aircraft designs. This report presents an investigation into the feasibility of developing a reconfigurable aircraft design system.

Firstly, the issues preventing current aircraft design systems from being used in a reconfigurable way are identified. These issues are closely tied to the intricate and tightly integrated nature of the source code that underpins these systems. This is unsurprising, given that these systems have typically been devised by experts in aircraft design (with varying expertise in software design) who are primarily interested in solving concrete design problems rather than creating sophisticated source code. In particular, the extensive design logic, characterized by high cyclomatic complexity, combined with cluttered and ambiguous design data structures, makes it challenging to comprehend and adapt the functioning of these systems.

An iterative development methodology is employed to come up with a software architecture aimed at mitigating the identified issues. A number of distinct architectural elements are devised that leverage a centralized semantic data management approach and a standardized interface for the formulation of modular analysis & sizing methods. Furthermore, a prototype aircraft design system that serves as a reference implementation for this architecture is developed. The prototype incorporates a graph database, a self-explanatory ontology (defining the semantics of the data stored in the database), and several abstract base classes for encapsulating the computation logic contained within typical analysis & sizing methods used during aircraft design studies. Concrete instances of these base classes are supposed to interact with the database through a well-defined endpoint interface, which includes extensive logging capabilities.

The prototype aircraft design system exhibits promising characteristics: The semantic data management approach facilitates the creation of a genuinely unambiguous and flexible data model. Simultaneously, it helps uncover inconsistencies and limitations in the employed analysis & sizing methods. Treating analysis & sizing methods as modular and nested instances of standardized classes appears to be the key to achieving reconfigurability. In addition to the unit-testability of these classes, the self-visualization features incorporated within can significantly enhance the comprehensibility and transparency of the analysis & sizing methods.

The architecture development, repository configuration, ontology formulation, and interface generation took a significant amount of time. Furthermore, some critical challenges surfaced, necessitating further investigation. Specifically, some of the employed analysis & sizing methods feature limitations and implicit assumptions that are required for a synthesis system but may need to be revised before being used in a reconfigurable design system based on the proposed architecture. In the end, there was insufficient time left for implementing a comprehensive set of analysis & sizing methods essential for materializing a thorough aircraft design loop. Therefore, achieving a fully functional aircraft design system prototype proved unattainable. Consequently, it was not possible to demonstrate that adopting the proposed architecture yields an aircraft design system that can be used in a reconfigurable way, despite indications that this could very well be the case...

The demand for environmentally friendly aviation has increased the interest in alternative concepts, such as hybrid-electric propulsion systems. A novel conceptual hybrid-electric sizing methodology was devised, which provides a detailed engine performance prediction and an accurate sizing of the powertrain components. The performance prediction methodology individually scales user-provided origin fan and engine core performance maps to generate a coherent engine performance map according to prescribed point performance requirements, while the component sizing methodology estimates individual component masses and determines their positions based on defined knowledge-rules. For each methodology, a dedicated design tool was created for easy integration into conceptual design workflows. A verification of the methodology via case studies assessing generated example aircraft concepts confirmed not only the validity of the obtained results, but also demonstrated a high accuracy at low computational costs. The exhibited robustness and sensitivity encourage the use of the methodology for future studies on hybrid-electric propulsion.

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

Zero-carbon-dioxide-emitting hydrogen-powered aircraft have, in recent decades, come back on the stage as promising protagonists in the fight against global warming. Nevertheless, most recent studies agree that hydrogen aircraft would underperform their kerosene counterparts in terms of operative empty mass and specific energy consumption. The main cause for the drop in performance lays in the fuel storage, as not only the liquid hydrogen has to be kept in cryogenic conditions and pressurised, but for the same energy content, it has four times the volume of kerosene. The inevitable consequences are an increase in fuselage size, which adds mass and drag to the aircraft, and the addition of an heavy fuel storage and distribution system. On the other side, hydrogen has 2.8 times higher specific energy, and the consequent reduction in fuel mass could balance the previously mentioned drawbacks. Literature on the topic shows that the optimal fuel storage solution depends on the aircraft mission, but most studies disagree on what solutions are optimal for each aircraft range category.The objective of this research was to identify and compare possible solutions to the integration of the hydrogen fuel containment system on short, medium and long-range airliners. The capabilities of an automated synthesis program for CS-25 aircraft have been expanded with validated structural and thermodynamic physics-based tank design models, to allow for the design and analysis of liquid hydrogen aircraft. Studies were performed on several design options. The effect of using an integral tank structure was found to be negligible for short-range aircraft, but increasingly more beneficial for medium and long-range aircraft. The effect of increasing the fuselage diameter was found to be favourable, especially when seats abreast could be added without the addition of one aisle. The effect of using a combination of an aft and a forward tank was found to be detrimental in terms of operational empty mass, beneficial in terms of specific energy consumption and negligible in terms of maximum take-off mass. The use of spherical tanks was found to be slightly beneficial, but only when compared to a non-spherical tank version using the same tank layout, non-integral tank structure, and same cabin layout. The study on the venting pressure revealed that with increasing aircraft size the optimal venting pressure in terms of main aircraft performance decreases whereas the sensitivity to those same parameters to the choice of venting pressure increases. The use of direct gas venting as a means to contain the pressure rise did not appear to provide significant performance improvements. The optimal designs, in terms of operational empty mass, maximum take-off mass and specific energy consumption, feature increased fuselage diameters, the use of the aft & forward tank layout, non-spherical tanks and no direct venting. The short-range aircraft uses non-integral tanks and high venting pressure, while the medium and the long-range aircraft benefit from an integral tank structure and a lower venting pressure. Nevertheless, the sensitivity to these design choices is not significant, meaning that with a different set of assumptions and/or requirements different design choices may become optimal. The overall best performing LH2 aircraft for the short, medium and long-range categories were found to have respectively 8%, 24% and 22% higher operative empty mass, -2%, 1% and -5% higher maximum take-off mass and 5%, 13% and 5% higher specific energy consumption than their kerosene versions.

A formulation based on the evaluation of power conversions within the system allows to overcome the limitations of conventional momentum analyses to study tightly coupled airframe-propulsion configurations such as Boundary Layer Ingestion. The so-called power balance method is applied in a CFD framework to gain insight about the application of this formulation for a wide range of flow conditions and for complex geometries. Not addressed in detail in previous research, it is quantified in this work the evolution of the distribution of energy in the wake of a body subject to turbulence, compressibility effects and angle-of-attack of the incoming flow. For the first time, a three-dimensional geometry representative of a typical aircraft is investigated with the power balance method. The results of a grid convergence study for the application of the power balance method in a CFD framework highlight the importance of generating a mesh with a fine resolution of the flow field globally, paying special attention to the development of the body wake.

The commercialization of the space industry has led to a reduction in size and weight of satellites and launching vehicles, which have effectively reduced the cost of space services. The costs of launching payloads to space can be significantly reduced when the launching vehicles are reusable. The two-stage-to-orbit system with a winged, reusable first stage vehicle, is deemed to offer benefits in terms of operations, as it can take-off and land from a runway.
Dawn Aerospace is pursuing development of the winged semi-Reusable Launching Vehicle with horizontal take-off and landing capabilities in the Mk-III concept. In the previous work on this concept, the aerodynamic design has not been addressed. It is important that the aerodynamic design is already considered in the conceptual design, as the large variations in flight conditions during the mission pose conflicting requirements on the aerodynamic design.
A performance analysis model of the launching vehicle was developed, in which the aerodynamic discipline is integrated. Using the developed model, sensitivity analyses and case studies were performed to investigate the impact of design changes on the mission performance. These results indicate that the limited gliding range of the first stage vehicle influences the propellant mass fraction of the upper stage.
The performed case studies indicate how the propellant mass fraction of the upper stage can be influenced, by changes to the vehicle design. Analysis of an alternative wing concept that improves the gliding performance of the first stage vehicle shows that the propellant mass fraction of the upper stage vehicle can be reduced from 91.5% to 90.3%. This can be achieved as the improved gliding range allows to reduce the delta-V delivered by the upper stage, and a less steep ascent trajectory that results in 8.9% less gravity losses. Analysis of a changed fuselage configuration indicates a reduction of propellant mass fraction from 91.5% to 90.4%.
Also changes to the trajectory design were analyzed, and by either reversing the direction of the take-off maneuver or by reserving propellant for the return trajectory, the propellant mass fraction can be reduced. The reductions achieved in the case studies were 0.5% for the reversed take-off direction, and 1.2% for the propelled return trajectory.