The economic case for hybrid-electric aircraft is often assessed on a single design mission, with less attention to airline network constraints and battery ageing. This paper presents a feedforward framework coupling conceptual aircraft sizing with a mixed-integer linear programming (MILP) tail-routing model that embeds battery degradation directly in the fleet operatingcost objective. Hybrid-electric variants of a regional turboprop (design ranges 600 to 1400 km, hybridisation 0 to 30%) are evaluated across two European networks and three economic scenarios. Under the reference cost assumptions, hybridising the reference aircraft raises cost per available seat kilometre (CASK) by about 17%, driven by utilisation, since charging lengthens turnarounds and requires an extra aircraft to serve the same demand. Modelling battery degradation changes the charging strategy: a fixed-coefficient degradation formulation selects schedules whose audited battery cost is 75% higher than self-assessed, yet the choice between plausible ageing scenarios reprices operations by only ±2%. Range margin, nearly free conventionally, costs up to 26% of CASK at high hybridisation, and under high fuel prices the cost-minimising design decouples from the minimum-weight design by 4 to 6%. Network operations and battery ageing can be first-order inputs to regional hybrid-electric aircraft design. ... Read more

Aviation was responsible for 2.5% of global energy-related CO2 emissions in 2023, and the market is projected to double over the next two decades. This growth conflicts with international targets to reach net-zero emissions by 2050. Of the available mitigation pathways, reducing in-flight energy consumption is the most immediately deployable, as it does not depend on new ground infrastructure. Aerodynamic drag is the main contributor to in-flight energy demand and is largely governed by the main wing. Since fixed wings are optimized for a single design point, they operate off-design during much of a mission. Morphing wings can adapt their shape in flight, offering a route to lower drag across a wider operating envelope.

This thesis assesses the effect of implementing chordwise two-degree-of-freedom (2-DoF) compliant camber-twist morphing trailing-edge wings on the sizing, configuration, and performance of CS-25 transport aircraft. Block fuel consumption is used as the primary performance metric, as it captures the trade-off between improved aerodynamic efficiency and the sizing penalties introduced by morphing systems.

Three methodological contributions support this assessment. First, a parameterization method was developed to describe the morphed shape of an arbitrary base airfoil using the rear spar location and two pseudo-deflection angles. Control points on the camberline are rotated and fitted with a fourth-order polynomial, after which the airfoil surfaces are reconstructed while preserving suction-side arc length to reflect compliant structure constraints. Applied at multiple spanwise stations, this method represents camber-twist morphing at wing level.

Second, the Aircraft Trimmed Performance Analysis Tool (ATPAT) was developed to evaluate trimmed cruise aerodynamic efficiency of morphing-wing aircraft. ATPAT combines a vortex lattice method with strip theory and sweep theory to resolve induced, viscous, pressure, and wave drag with runtimes of seconds to one minute per evaluation. The tool was validated against experimental and computational data on the Fokker 100 and integrated into the Initiator aircraft design toolbox with Bayesian and gradient-based optimizers.

Third, two uncertain sizing inputs were explicitly addressed: morphing system mass and achievable high-lift increment. Morphing system mass was treated as a parameter bounded by the specific mass of Fowler flaps, while RANS CFD estimated a conservative sectional lift increase of 0.63 at full deflection.

The methodology was applied to the ATR 72-600 and Fokker 100. For both aircraft, morphing designs were heavier and required larger wing areas than conventional designs due to high-lift limitations, making field performance the main constraint. Relaxing landing field length by about 20% brought sizing outcomes closer to the reference aircraft.

For the ATR 72-600, aerodynamic gains did not offset the penalties. The morphing design matched conventional block fuel only on the harmonic mission and used 6% and 4% more fuel on shorter and maximum-fuel missions. For the Fokker 100, a block-fuel reduction of up to 3.6% was achieved on the harmonic mission after a 17.5% relaxation of landing field length, with sizing nearly identical to the reference aircraft. Its transonic cruise also offers additional drag reduction potential through wave drag reduction.

The results remain subject to uncertainty. The relative contributions of sizing effects and wave drag reduction were not isolated, morphing system mass remains uncertain, and high-lift performance estimates are based on limited airfoil data. More detailed experimental validation and segmented mission analysis are recommended to refine the estimated benefits of morphing wings.
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Engineering design frequently involves optimisation problems incorporating computationally expensive analysis tools. Surrogate-based optimisation has been shown to reduce the runtime of these optimisations. Multi-point (MP) and multi-fidelity (MF) infill strategies can further accelerate convergence, yet their combination has not been studied extensively.
This thesis introduces and evaluates an asynchronous MP MF infill strategy, benchmarked against eleven unconstrained MF numerical problems, using expected runtime (ERT).
The asynchronous MP single-fidelity strategy (16 ranks) reduced the geometric mean ERT by 72.7% compared with the baseline Efficient Global Optimisation with Expected Improvement. The single-point MF strategy, on the other hand, increased the ERT by 21.2%, degrading performance. The combined asynchronous MP MF strategy achieved a 68.2% reduction relative to the baseline.
These results show that the asynchronous MP strategy substantially improves performance. In contrast, the selected MF strategy proves detrimental, indicating that a revised MF strategy is required to yield further gains. ... Read more

Distributed electric propulsion (DEP) modifies the spanwise lift distribution through propeller slipstream interaction, directly affecting wing-root bending loads and structural sizing. This paper extends a mid-fidelity aero-propulsive–structural framework to evaluate wing structural sensitivity for distributed-propulsion hybrid-electric regional aircraft with multiple propellers at arbitrary spanwise locations. Blade pitch trimming under non-uniform inflow is incorporated to obtain physically consistent slipstream velocities, and a three-dimensional numerical vortex-image correction replaces the previously used analytical jet correction to remove geometric constraints and enable flexible propulsion layouts. The objective is to quantify which propulsion-integration parameters govern wing structural mass and to assess whether clean-wing structural sizing assumptions remain valid for DEP configurations. The framework is applied to a representative 70-passenger hybrid-electric regional aircraft under cruise and 2.5𝑔 pull-up conditions. Results show that wing mass is governed primarily by spanwise load redistribution and associated bending-moment arm effects. Parameters that shift lift outboard are structurally penalising, while inboard redistribution reduces mass. Propeller rotational direction and thrust-share allocation emerge as first-order drivers, producing mass variations up to ∼ 3%, whereas swirl magnitude, spanwise translation, and disk-area changes are secondary within the investigated design space. Clean-wing aerodynamic loads are therefore not universally conservative and can underpredict structural mass depending on thrust distribution and propulsor placement. These findings demonstrate that propulsion-integration decisions must be evaluated jointly with structural sizing in conceptual DEP aircraft design ... Read more

Liquid hydrogen-powered aircraft are a promising candidate for reducing aviation’s climate impact, but safe integration requires crashworthiness considerations already at conceptual design. This work presents a parametric, knowledge-based methodology to design and analyse a fuselage-located cryogenic tank supported by a belly crash structure under a vertical drop test. The methodology is implemented in a parametric application that automatically generates geometry from readily adjustable inputs, enabling quick generation of a wide variety of configurations. A baseline configuration, consisting of 12 sub-tank X-beams per bay configured in six X-crosses, a stiff crossbeam design, and distributed Kevlar tank suspension ropes, exhibited favourable deformation kinematics. A design-of-experiments study was conducted and combined with a surrogate-model-based sensitivity analysis to identify the dominant design parameters governing the crash response. Results show that the coupled frame and crossbeam thickness, together with the crash coefficient, are the most influential. Including a crash structure substantially increases the required tank and fuselage length, and the associated mass penalty is dominated by the added tank and fuselage structure rather than by the crash structure mass. The developed model provides a basis for further research on structural mass optimisation and for providing data for a surrogate integration into a conceptual aircraft design tool. ... Read more

Hydrogen-fueled aircraft promise to achieve significantly reduced climate impact compared to their kerosene counterparts. However, while hydrogen has a significantly higher gravimetric density than kerosene, its volumetric density is much lower, meaning that even in a liquid state, it takes up approximately four times the volume. This results in significant challenges in integrating hydrogen storage in aircraft. It has been suggested that blended wing body (BWB) aircraft may provide a better-suited alternative for hydrogen integration compared to the conventional tube-and-wing (TAW) configuration.
Prior research has investigated hydrogen-fueled BWB aircraft, highlighting different potential tank integration strategies. To the author's knowledge, no direct comparison of these hydrogen configurations highlights their relative performance impact compared to a consistent kerosene baseline aircraft. Therefore, this work compares different hydrogen tank configurations identified in prior research to a kerosene BWB baseline under consistent top-level aircraft requirements (TLARs), including a design range of 2500 nmi and 239 passengers. Technology assumptions for a 2050 entry into service (EIS) are applied. The analysis is complemented through the development of hydrogen and kerosene TAW configurations that fulfill the same TLARs.
The comparison is performed in three steps. First, a BWB is designed for each tank configuration using conceptual design methods. Second, the four tank layouts are compared and the layout is identified that imposes the lowest energy penalty on the BWB. Third, the integration penalty of the hydrogen BWB is compared to that of the hydrogen TAW.
A BWB employing a combined tank configuration, with hydrogen tanks located beside and aft of the passenger cabin, is found to experience the lowest integration penalty. The BWB experiences a 12.5% penalty in block energy for the design mission, compared to a penalty of 10.5% experienced by the TAW. This indicates that the BWB is less well suited to hydrogen integration than the TAW under the specified TLARs. In addition, sensitivity studies are conducted to evaluate the impact of integration assumptions on the BWB. These show that reasonable variations in the assumptions do not change the conclusion of this study. The findings do not eliminate hydrogen-fueled BWB aircraft as a viable alternative to hydrogen-fueled TAW designs. In fact, literature shows that BWB configurations still offer an inherent efficiency advantage, although their higher integration penalty must be considered in the overall trade-off. ... Read more

This work presents the methodology for the conceptual modeling and preliminary sizing of Solid Oxide Fuel Cell–Gas Turbine (SOFC–GT) hybrid powertrains within a conceptual aircraft design framework. SOFC–GT hybrid systems, long investigated in commercial power generation, offer higher combined power output, improved overall thermal efficiency, and greater fuel flexibility compared to similar FC– GT architectures. As high-temperature solid oxide fuel cells (HT–SOFCs) are not electrochemically bound by a theoretical efficiency limit like the Carnot cycle for gas turbines, their integration with conventional engines has the potential to significantly improve system efficiencies while retaining high power densities. The primary objective of this work is to evaluate their integration into regional turboprop aircraft and compare them against conventional aircraft with kerosene/SAF and LH2-based gas turbine powertrains for similar Top Level Aircraft Requirements (TLAR). The sizing methodologies for SOFC-GT aircraft are incorporated within the Class 1 sizing loop of the Aircraft Design Initiator (ADI), an in-house conceptual aircraft design tool developed by the Faculty of Aerospace Engineering at Delft University of Technology for rapid design evaluation. For this, the powertrain model previously developed by De Vries (for hybridelectric powertrains) and Borgia (for hydrogen-based hybrid-electric powertrains) was extended for the sizing of SOFC-GT aircraft. The modeling approach starts at the fuel cell level, using validated electrochemical and thermodynamic models for the SOFC. A baseline tubular SOFC geometry is selected and modified to improve stacklevel gravimetric power density through reductions in component thicknesses, dominant ohmic losses, and an increase in the effective cell area. The cell-level model is then integrated into a system-level representation of the SOFC-GT powertrain, which is then incorporated into the ADI’s Class 1 sizing loop to meet the propulsion and power requirements of the baseline ATR 72-600 aircraft. At the system level, the Class 1 sizing results of the SOFC-GT aircraft (at a design current density of 400 mA/cm², cruise H2 split of 50%, and 75% fuel utilization) show an overall thermal efficiency of 44.17% in cruise, representing a 48% improvement over the baseline ATR 72-600 and 44% over the LH2-based aircraft. These gains are accompanied by increases in MTOM (19.2% and 14.7%) and OEM (+45.2% and +23%) compared to kerosene and LH2 aircraft, respectively. Nevertheless, the higher efficiency reduces the required energy mass for the full-range design mission by 65% and 13.5%, while the larger wingspan (29.44 m) remains within ICAO Class C limits. Consequently, the lower H2 demand also decreases H2O and NOX emissions by 15.3% and 50.26% relative to conventional LH2 aircraft. Results from parametric analyses indicate that the hydrogen power split between the gas turbine and the SOFC stacks in cruise is the most influential design parameter for maximizing performance, with higher values improving overall thermal efficiency but also significantly increasing the aircraft operating empty mass (OEM) and energy requirements for the design mission. The trade-off between fuel cell and gas turbine contributions is sensitive to fuel utilisation, where low utilisation reduces efficiency by shifting power production toward the less efficient gas turbine. On the other hand, increasing anode fuel utilisation above 75% improves overall thermal efficiency but can reduce gravimetric power density at the stack level due to mass transport losses, the effect of which can be partially mitigated by increasing the excess air supply at the cathode. Additionally, the SOFC system pressure and operating temperature were also observed to be important design parameters that could improve the performance of the SOFC-GT powertrain. However, the cruise thermal efficiency exhibits an asymptotic trend at higher pressures and temperatures, beyond which the benefits diminish. At elevated temperatures, further gains are offset by an increase in system mass, while at higher pressures, the additional Balance of Plant (BoP) power requirements lead to reductions in both gas turbine and SOFC efficiencies. ... Read more

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. ... Read more

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

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. ... Read more

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. ... Read more

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. ... Read more

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. ... Read more

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. ... Read more

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. ... Read more

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... ... Read more

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. ... Read more

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. ... Read more

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. ... Read more