The feasibility of converting an existing internal combustion powered general aviation aircraft to battery electric propulsion was studied. The theoretical performance of various types of airframes with battery electric propulsion systems was compared to determine which type of airframe would be best suited for conversion. It was found that battery electric propulsion is best used in aircraft intended for slow speed, efficient flight and carrying limited payload which is a mission typically flown in motor gliders. A reference motor glider was selected and a conceptual power system packaging design study was performed. The study determined that a critical component of the power system packaging design was the technical feasibility of packaging the batteries inside of the wing structure. This was driven by center of gravity restrictions. Technical concerns related to a battery-in-wing design were investigated, included wing aeroelastic performance, wing stiffness and wing strength. The results showed that aeroelastic flutter was not a driving design criteria for the reference airframe used as the physical size of the battery did not allow for them to be packaged in wing locations that detrimentally affected flutter performance. The battery packaging layout was instead driven by access for battery maintenance, battery safety and the battery thermal management system. Overall weight change from packaging the batteries in the wing compared to the fuselage was found to be negligible. The resulting aircraft conceptual design indicated a powered flight range with reserves of over 200 miles and a powered flight endurance of greater than 3 hours with 2 persons onboard.

In traditional aircraft conceptual design, subsystems are largely accounted for implicitly based on available historical data and trends. Such an approach has limitations when novel subsystem architectures such as More Electric or All Electric aircraft are considered, since historical data regarding such architectures is either limited or non-existent. In such cases, the incorporation of more thorough and explicit consideration of the aircraft subsystems into the conceptual design phase is warranted. The first objective of this dissertation is to integrate subsystem sizing and analysis methods that are suitable for the early design phases with the traditional aircraft sizing methodology. The goal is to facilitate the assessment subsystem architecture performance with respect to vehicle and mission level metrics. The second objective is to investigate how the performance of different subsystem architectures varies with aircraft size. The third and final objective is to assess the sensitivity of architecture performance to epistemic and technological uncertainty. These objectives are pursued through the development of an integrated sizing and analysis environment where the subsystems are sized in parallel with the aircraft itself using subsystem models that are computationally inexpensive and do not require detailed aircraft definition. The effects of subsystem mass, secondary power requirements, and drag increments are propagated to the mission performance analysis following which the vehicle and subsystems are re-sized. A number of experiments are performed to first test the capabilities of the developed environment and subsequently assess the performance of numerous subsystem architectures and the sensitivity of select architectures to epistemic and technological uncertainty.

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The aerospace industry has always been looking forward new developments with the aim to create more environmental friendly aircraft, as well as to improve their performance. Over the last few years, a prominent research topic to achieve these challenging goals has been focussed on the incorporation of more electric Secondary Power Systems (SPS), this concept is known as More Electric Aircraft (MEA) or All Electric Aircraft (AEA) when the internal combustion engine is also replaced. Among others, Airbus is using Electro-hydrostatic Actuators, (EHAs) to combine hydraulic and electric power in A320 and A340 for flight tests since 1993. The company TTTECH applied the same concept by working on the development of an electrical steering system for an aircraft nose landing gear, and power source rationalization and electrical power flexibility in aircraft. Some of the advantages stated when the MEA concept is applied are: reduction in aircraft weight and performance penalties related to conventional SPS. Although the More/All electric aircraft concept provided satisfactory results for fixed-wing aircraft, research for rotary-wing aircraft is less common. This encourages the assessment of fuel consumption and performance penalties due to conventional and more electric SPS at conceptual level, which could achieve similar outcomes, while finding the best configuration possible. This project takes into account the previous research focused on fixed-wing aircraft and studies on new technologies for SPS within Cranfield University, this includes electrical Ice Protection System (IPS), Environmental Control System (ECS) and Actuation System (AS). Additionally, Fuel System (FS) and Electrical System (ES) capabilities were added, developing a generic tool able to predict the total power requirements depending on the flight conditions. This generic tool was then integrated with a performance model, where overall fuel consumption is calculated for a flight mission, giving continuity and improvement to the work already done. Secondary systems configuration and operating characteristics for a representative light single-engine rotary-wing aircraft were tailored, and the systems behaviour is presented. Finally, fuel consumption was calculated for a baseline mission profile, and compared to the fuel consumption when the systems are not included. The baseline mission set the initial flight conditions from which a parametric study was carried out; by varying these conditions the parametric study determined total fuel requirements for the analysed flight segments. An increment of up to %1.9 in the fuel consumption was found by integrating the proposed systems to the performance model, showing the impact produced by the systems, and the importance of studying different technologies to minimise it.

Several models of battery electric aircraft are currently under development and are expected to be commercialized in the coming years. These electric aircraft are expected to require a significant amount of electrical power within a short period of time. Sweden is currently facing local and regional capacity issues in its electrical grid, leading to a decline in grid connection applications from new businesses. The objective of this thesis is to investigate the necessary charging infrastructure on-site to support an electrified air fleet. Additionally, the thesis aims to explore how local battery storage systems and a PV plant could be utilized to ensure a robust and resilient operation. In order to address these questions, a model of an airport has been developed using Matlab Simulink. A case study was conducted for an airport to demonstrate the practical application of the model. The results indicate that charging battery electric aircraft will impose a volatile power demand on the grid, with high peaks. Inadequate power supply leads to queuing issues and longer turnaround times. The results also highlight the significance of a battery storage system, as it enables the handling of more aircraft. A PV plant complements the battery storage system well, as it produces the majority of its power during peak traffic hours. The simulations further demonstrate that the PV plant helps recharge the battery storage system between the morning and afternoon traffic peaks. Moreover, detailed aircraft arrival data for the investigated airport is crucial for obtaining accurate results.

Envisioned in the next 15 to 30 years in the aviation industry, hybrid-electric propulsion offers theopportunity to integrate new technology bricks providing additional degrees of freedom to improveoverall aircraft performance, limit the use of non-renewable fossil resources and reduce the aircraftenvironmental footprint. Today, hybrid-electric technology has mainly been applied to groundbased transports, cars, buses and trains, but also ships. The feasibility in the air industry has to beestablished and the improvement in aircraft performance has still to be demonstrated. This thesisaims to evaluate the energy savings enabled by electric power in the case of a 70-seat regionalaircraft. First, energy saving opportunities are identified from the analysis of the propulsion andaerodynamic efficiencies of a conventional twin turboprop aircraft. The potential benefits comingfrom the variation of the size of prime movers and the new power managements with the use ofbatteries are studied. Also, possible aerodynamic improvements enabled by new propellerintegrations are considered. For each topic, simplified analyses provide estimated potential ofenergy saving. These results are then used to select four electrified propulsion systems that arestudied in more detail in the thesis: a parallel-hybrid, a turboelectric with distributed propulsion, apartial-turboelectric with high-lift propellers and an all-electric. Evaluating the selected hybrid-electric aircraft is even more challenging that the sizing of the different components, the energymanagement strategies and the mission profiles one can imagine are many and varied. Inaddition, the overall aircraft design process and the evaluation tools need to be adaptedaccordingly. The Airbus in-house Multidisciplinary Design Optimisation platform named XMDO,which includes most of the required modifications, is eventually selected and further developedduring the thesis. For examples, new parametric component models (blown wing, electrical motor,gas turbine, propeller, etc…) are created, a generic formulation for solving the propulsion systemequilibrium is implemented, and simulation models for take-off and landing are improved. In orderto evaluate the energy efficiency of the hybrid-electric aircraft, a reference aircraft equipped with aconventional propulsion system is first optimised with XMDO. Different optimisation algorithms aretested, and the consistency of the new design method is checked. Then, all the hybrid-electricconfigurations are optimised under the same aircraft design requirements as the reference. Forthe electrical components, two levels of technology are defined regarding the service entry date ofthe aircraft. The optimisation results for the turboelectric and the partial-turboelectric are used tobetter understand the potential aerodynamic improvements identified in the first part of the thesis.Optimisations for the parallel-hybrid, including different battery recharge scenarios, highlight thebest energy management strategies when batteries are used as secondary energy sources. All theresults are finally compared to the reference in terms of fuel and energy efficiencies, for the twoelectrical technology levels. The last part of the thesis focuses on the all-electric aircraft, and aimsat identifying the minimum specific energy required for batteries as a function of the aircraft designrange. A trade study is also carried-out in accordance with the service entry date for the otherelectrical components

L'avion plus électrique est un concept qui a le vent en poupe chez les principaux constructeurs du domaine de l'aéronautique. Dans ce domaine, comme dans d'autres, les besoins en énergie électrique sont croissants et nécessitent de mettre en place des systèmes de conversion d'énergie fiables, performants et modulaires. Ces systèmes de conversion sont souvent couplés avec des systèmes de stockage d'énergie (type batterie) permettant dans certaines situations de rendre l'avion énergétiquement autonome grâce à une source de puissance indépendante des principaux organes de production d'énergie. Cette interconnexion batterie - réseau de bord présente un rapport de tension élevé ce qui, ajouté aux fortes valeurs de courant de la basse tension, en fait un objet particulièrement complexe à réaliser.L'objectif de cette thèse est de concevoir de manière optimale un convertisseur de puissance isolé permettant l'interconnexion d'un bus basse tension 28 V (typiquement des batteries) à un bus haute tension 540 V (réseau de bord de l'avion) avec une puissance échangeable d'environ 12 kW. Elle se déroule dans le cadre d'un projet ANR (quatre partenaires universitaires, associés à l'entreprise AIRBUS) dont l'une des tâches est le développement d'outils de conception pour l'électronique de puissance. Le travail correspondant comprend une contribution à cette tâche sous forme de la construction de modèles des principaux composants intervenant dans un convertisseur, modèles destinés à être intégrés dans les routines d'optimisation. Pour cette raison, ils seront analytiques (physique, empiriques, mélange des deux).Ces modèles seront ensuite insérés dans un outil global développé dans une autre thèse du projet, à l'aide duquel différentes architectures de convertisseurs seront comparées afin d'en déduire la meilleure solution pour le cahier des charges énoncé précédemment. Un prototype du convertisseur retenu sera finalement réalisé en utilisant des technologies avancées, pour conduire une validation expérimentale. / The electric aircraft tends to become widespread at all the main manufacturers of the domain of the aeronautics. Needs do not stop growing and require setting up reliable, efficiency and modular systems of conversion of energy. These systems of conversion are often coupled with systems of storage of energy (battery) allowing in certain situations to make the punctually autonomous aircraft energetically thanks to a source of power independent from main organs of power production. This interconnection battery - network of edge presents a very high report of rise of tension what, added to the high current value of the battery bus, in fact a particularly complex object to be realized.The objective of this thesis is to design in an optimal way a converter of power isolated allowing the interconnection of a low-voltage bus 28V (typically batteries) in a high-voltage bus 540V (network of edge of the aircraft) with an exchangeable power about 12 kW. It takes place within the framework of an ANR project (four university partners + AIRBUS) the development of tools of conception of which one of the tasks is for the ENP. The corresponding work includes a contribution to this task in the form of the construction of models of the main components occurring in a converter, model intended to be integrated into the routines of optimization. For that reason, they will be analytical (physical, empirical or mix both).These models will then be inserted into a global tool developed in another thesis of the project, by means of which various architectures of converters will be compared to deduct the best solution from it for the previous specifications. A prototype of the reserved converter will be finally realized by using advanced technologies, to lead an experimental validation.

This thesis presents the design and analysis of specialised control systems for switched reluctance (SR) and permanent magnet (PM) machines in vehicular electric applications. Control systems for operation in motoring and power generation are considered for both the types of machines. The SR machine operation considered in this thesis is mainly focused towards the application of aero-engine starter/generators. The control designs for PM machines are formulated considering general fault-tolerant and isolated multiphase PM machines which can be applied in the majority of safety-critical vehicular power and propulsion applications. The SR motoring mode presented in this thesis considers the control design for operation from zero speed to a high speed range, while SR generation mode is confined to the high speed range, such as for the requirements of aero-engine starter/generator operation. This thesis investigates applied control methods for both single-pulse and chopping modes of operation. Classical excitation control versus peak current control and the introduction of a zero-voltage interval are compared for SR motor operation. Optimized excitation control versus two classical forms of excitation control are developed and compared for SR generator operation. Studies include simulation of a 12/8 250kW machine and experimental work on a 6/4 300W machine. The PM motoring and power generation considered in this thesis focuses on a special class of PM machines and drives which are specifically designed for fault-tolerant operation. Optimized control strategies for the operation of PM machines with the parallel H-bridge per-phase converter architecture are investigated. Mathematical modelling of the machine and drive with a consideration of harmonics is presented. The developed control methods are then evaluated by means of finite-element model based simulations of a 125kW five phase surface PM rotor machine and an interior PM rotor machine.

Modern 'more-electric' aircraft demand increased levels of electrical power as non-propulsive power systems are replaced with electrical equivalents. This electrical power is provided by electrical generators, driven via a mechanical transmission system, from a rotating spool in the gas turbine core. A wide range of electrical loads exist throughout the aircraft, which may be pulsating and high powered, and this electrical power demand is transferred though the generators to produce a torque load on the drivetrain. The mechanical components of the drivetrain are designed for minimum mass and so are susceptible to fatigue, therefore the electrical loading existing on modern airframes may induce fatigue in key mechanical components and excite system resonances in both mechanical and electrical domains. This electro-mechanical interaction could lead to a reduced lifespan for mechanical components and electrical network instability.This project investigates electro-mechanical interaction in the electrical power offtake from large diameter aero gas turbines. High fidelity modelling of the drivetrain, and generator, allow the prediction of system resonances for a generic gas turbine-generator system. A Doubly-Fed Induction Generator (DFIG) is considered and modelled. DFIGs offer opportunities due to their fast dynamics and their ability to decouple electrical and mechanical frequencies (e.g. enabling a constant frequency electrical system with a variable speed mechanical drive). A test platform is produced which is representative of a large diameter gas turbine and reproduces the electro-mechanical system behaviour. The test platform is scaled with respect to speed and power but maintains realistic sizing between component dimensions which include: a gas turbine mechanical spool emulation, transmission driveshafts and gearbox, and accessory loads such as a generator. This test platform is used to validate theoretical understanding and suggest alternative mechanical configurations, and generator control schemes, for the mitigation of electro-mechanical interaction.The novel use of a DFIG and an understanding of electro-mechanical interaction allow future aircraft designs to benefit from the increased electrification of systems by ensuring that sufficient electrical power can be provided by a robust gas turbine-generator system.

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