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The purpose of this paper is to analyze the electric propulsion use in civil aviation and propose a framework for certification of electric propulsion subsystems. Although electric propulsion architectures are discussed as key technology for the future of aviation, the industry standards as well as regulations fail to cover the application in full extent, specifically for commercial large airplanes. This paper proposes an approach for the analyses of reliability and certification of the new-generation propulsion system by pointing out the “common structure” among the possible architectures.

The research process used in this paper consists of following steps: the challenges of the hybrid-electric propulsion is listed, the architectures of the hybrid-electric applications in the literature are identified, the differences of the hybrid architectures from the present applications by means of application and standardization are discovered, the architectures are analyzed and the two main subsystems are defined – the present combustion system and the common unit, which is a similar structure used in all-electric aircraft. For this purpose, the standards used for design basis and certification of the present propulsion system and their relationship with the subsystems of the architectures have been analyzed. The procedure for the reliability assessment of the system is given, a framework for the safety assessment and the certification of the propulsion systems is proposed to make it easier and without sacrificing the already accumulated experience. This study shows that by using the common unit, the present certification framework can be used, by focusing on the reliability of the common unit and its integration with the rest of the architecture.

A specific definition of common unit is proposed, to point out the difference in certification efforts of hybrid-electric propulsion architectures. Yet, there is no data available for propulsion-level airborne battery and electrical systems to assess the reliability. Thus, dividing the propulsion system into two main systems and providing a model for certification of the common unit sub-system would be beneficial for easy deployment of the hybrid architectures both for design and for certification. In this paper, it is proposed that by using this common unit, the present certification framework can be used as it is, by focusing on the reliability of the common unit and its integration with the rest of the architecture.

The aircraft certification regulations act in two ways: they provide a starting point for new design projects, and they are a basis for certification of the final system. This study aims to draw focus on certification issues on the new-generation hybrid-electric propulsion systems. With the introduction of hybrid-electric propulsion for large aircraft, the present standards (CS-25, CS-E, CS-P, CS-Battery and CS-APU) create an obstacle for further progress as their borders get into each other. Instead of developing a new set of standard(s), this paper proposes a new approach by dividing the propulsion system into two subsystems.

This research proposes a definition of “common unit” for simplification of the hybrid-electric propulsion architectures for large civil aircraft. The common unit consists of both battery and electrical components and their reliability shall be considered for hybrid-electric propulsion.

This paper aims to present the main results achieved in the frame of the TIVANO national-funded project which may anticipate, in a stepped approach, the evolution and the design of the enabling technologies needed for a hybrid/electric medium altitude long endurance (MALE) unmanned aerial vehicle (UAV) to perform persistent intelligence surveillance reconnaissance (ISR) military operations.

Different architectures of hybrid-propulsion system are analyzed pointing out their operating modes to select the more suitable architecture for the reference aircraft. The selected architecture is further analyzed together with its electric power plant branch focusing on electric system architecture and the selected electric machine. A final comparison between the hybrid and standard propulsion is given at aircraft level.

The use of hybrid propulsion may lead to a reduction of the total aircraft mass and an increase in safety level. However, this result comes together with a reduced performance in climb phase.

This study can be used as a reference for similar studies and it provides a detailed description of propulsion operating modes, power management, electric system and machine architecture.

This study presents a novel application of hybrid propulsion focusing on a three tons class MALE UAV for ISR missions. It provides new operating modes of the propulsion system and a detailed electric architecture of its powertrain branch and machine. Some considerations on noise emissions and infra-red traceability of this propulsion, at aircraft level.

The aim of this paper is to assess the potential of fuel-battery hybrid narrow-body (180PAX) transport aircraft according to different design ranges for an entry-into-service (EIS) of 2035.

The philosophy used in the design of the twin-engine fuel-battery hybrid concept is to use the power of an electric motor during cruise to drive a single propulsive device, whereas the other one is powered conventionally by an advanced gas turbine. A methodology for the sizing and performance assessment of hybrid energy aircraft was previously proposed by the authors. Based on this methodology, the overall sizing effects at aircraft level are considered to size the hybrid aircraft to different range applications. To evaluate the hybrid concept, performance was contrasted against a conventional aircraft projected to EIS 2035 and sized for identical requirements. Additionally, sensitivity of the prospects against different battery technology states was analysed.

The best suited aircraft market for the application of the fuel-battery hybrid transport aircraft concept considered is the regional segment. Under the assumption of a battery-specific energy of 1.5 kWh/kg, block fuel reduction up to 20 per cent could be achieved concurrently with a gate-to-gate neutral energy consumption compared to an advanced gas-turbine aircraft. However, a large increase in maximum take-off weight (MTOW) occurs resulting from battery weight, the additional electrical system weight, and the cascading sizing effects. It strongly counteracts the benefit of the hybrid-electric propulsion technology used in this concept for lower battery-specific energy and for longer design ranges.

The findings will contribute to the evaluation of the feasibility and impact of hybrid energy transport aircraft as potential key enablers of the European and US aeronautical program goals towards 2035.

The paper draws its value from the consideration of the overall sizing effects at aircraft level and in particular the impact of the hybrid-electric propulsion system to investigate the prospects of fuel-battery hybrid narrow-body transport aircraft sized at different design ranges.

When comparing and contrasting different types of fixed-wing military aircraft on the basis of an energetic efficiency figure-of-merit, unmanned aerial vehicles (UAVs) dedicated to tactical medium-altitude long-endurance (MALE) operations appear to have significant potential when hybrid-electric propulsion and power systems (HEPPS) are implemented. Beginning with a baseline Eulair drone, this paper aims to examine the feasibility of retro-fitting with an Autarkic-Parallel-HEPPS architecture to enhance performance of the original single diesel engine.

In view of the low gravimetric specific energy performance attributes of batteries in the foreseeable future, the best approach was found to be one in which the Parallel-HEPPS architecture has the thermal engine augmented by an organic rankine cycle (ORC). For this study, with the outer mould lines fixed, the goal was to increase endurance without increasing the Eulair drone maximum take-off weight beyond an upper limit of +10%. The intent was to also retain take-off distance and climb performance or, where possible, improve upon these aspects. Therefore, as the focus of the work was on power scheduling, two primary control variables were identified as degree-of-hybridisation for useful power and cut-off altitude during the en route climb phase. Quasi-static methods were used for technical sub-space modelling, and these modules were linked into a constrained optimisation algorithm.

Results showed that an Autarkic-Parallel-HEPPS architecture comprising an ORC thermal energy recovery apparatus and high-end year-2020 battery, the endurance of the considered aircraft could be increased by 11%, i.e. a total of around 28 h, including de-icing system, in-flight recharge and emergency aircraft recovery capabilities. The same aircraft with the de-icing functionality removed resulted in a 20% increase in maximum endurance to 30 h.

Although the adoption of Series/Parallel-HEPPS only solutions do tend to generate questionable improvements in UAV operational performance, combinations of HEPPS with energy recovery machines that use, for example, an ORC, were found to have merit. Furthermore, such architectural solutions could also offer opportunity to facilitate additional functions like de-icing and emergency aircraft recovery during engine failure, which is either not available for UAVs today or prove to be prohibitive in terms of operational performance attributes when implemented using a conventional PPS approach.

This technical paper highlights a new degree of freedom in terms of power scheduling during climbing transversal flight operations. A control parameter of cut-off altitude for all types of HEPPS-based aircraft should be introduced into the technical decision-making/optimisation/analysis scheme and is seen to be a fundamental aspect when conducting trade-studies with respect to degree-of-hybridisation for useful power.

Velomobiles or bicycles cars are human-powered vehicles, enclosed for improving aerodynamic performance and protection from weather and collisions. The purpose of this paper is to design and develop a three-wheeled velomobile equipped with a hybrid human-electric propulsion system.

The mechanical layout has been developed in order to improve safety, a CAD code has been used for the design and the dynamic performances have been studied by means of specific multi-body codes. The electric propulsion system has been designed both with analytical and FEM methods.

A special three-wheeled tilting vehicle layout equipped with a four-bar linkage connection has been developed. A particular synchronous reluctance machine has been developed, which is very suitable for human-electric hybrid propulsion. A MATLAB code for integrated mechanical and electrical analysis has been developed.

A new kind of light vehicle has been conceived. A new synchronous reluctance machine with high efficiency has been developed. A performance analysis in electric, human and hybrid working modes has been presented, which takes into account the specific features of both the electric motor and the pedaling legs. A prototype of the vehicle has been built.

A hybrid-electric unmanned aerial vehicle (HE-UAV) model has been developed to address the problem of low endurance of a small electric UAV. Electric-powered UAVs are not capable of achieving a high range and endurance due to the low energy density of its batteries. Alternatively, conventional UAVs (cUAVs) using fuel with an internal combustion engine (ICE) produces more noise and thermal signatures which is undesirable, especially if the air vehicle is required to patrol at low altitudes and remain undetected by ground patrols. This paper aims to investigate the impact of implementing hybrid propulsion technology to improve on the endurance of the UAV (based on a 13.6 kg UAV).

A HE-UAV model is developed to analyze the fuel consumption of the UAV for given mission profiles which were then compared to a cUAV. Although, this UAV size was used as reference case study, it can potentially be used to analyze the fuel consumption of any fixed wing UAV of similar take-off weight. The model was developed in a Matlab-Simulink environment using Simulink built-in functionalities, including all the subsystem of the hybrid powertrain. That is, the ICE, electric motor, battery, DC-DC converter, fuel system and propeller system as well as the aerodynamic system of the UAV. In addition, a ruled-based supervisory controlled strategy was implemented to characterize the split between the two propulsive components (ICE and electric motor) during the UAV mission. Finally, an electrification scheme was implemented to account for the hybridization of the UAV during certain stages of flight. The electrification scheme was then varied by changing the time duration of the UAV during certain stages of flight.

Based on simulation, it was observed a HE-UAV could achieve a fuel saving of 33% compared to the cUAV. A validation study showed a predicted improved fuel consumption of 9.5% for the Aerosonde UAV.

The novelty of this work comes with the implementation of a rule-based supervisory controller to characterize the split between the two propulsive components during the UAV mission. Also, the model was created by considering steady flight during cruise, but not during the climb and descend segment of the mission.

Parallel hybrid electric (HE) propulsion retrofit is a promising alternative to reduce fuel burn of aircraft operating on short regional flights. However, if the batteries are depleted at the end of the mission, the hybrid powertrain designs with downsized gas turbines (GTs) and additional electric motors might not meet the one-engine inoperative (OEI) missed approach climb performance required by the certification. Alternatively, hybrid designs using the original full-size GT can perform one engine climb without electric assistance. This paper aims to evaluate the impact of overshoot climb requirements on powertrain design and performance comparing the two design approaches.

An aircraft-level parametric mission analysis model is used to evaluate aircraft performance combined with an optimization framework including multiple constraints. An indirect approach using metamodels is used to optimize powertrain sizing and operation strategy.

Considering OEI climb requirements, no benefits were found using a design with downsized GTs. Equivalent fuel burns were found for hybrid designs that keep the original size GTs, but do not require electric energy for the OEI overshoot at the end of the mission. Then, it is recommended to size the GT to maintain the emergency climb capabilities with no electric assistance to ensure power availability regardless of remaining battery energy.

This work introduces a new perspective on parallel HE sizing with consideration for the dependency of power capability at aircraft level on the electric energy availability in case of critical mission scenarios such as overshoot climb at the end of the mission.

The purpose of this study is to examine state-of-the-art in hybrid-electric propulsion system modeling and suggest new methodologies for sizing such advanced concepts. Many entities are involved in the modelling and design of hybrid electric aircraft; however, the highly multidisciplinary nature of the problem means that most tools focus heavily on one discipline and over simplify others to keep the analysis reasonable in scope. Correctly sizing a hybrid-electric system requires knowledge of aircraft and engine performance along with a working knowledge of electrical and energy storage systems. The difficulty is compounded by the multi-timescale dynamic nature of the problem. Furthermore, the choice of energy management in a hybrid electric system presents multiple degrees of freedom, which means the aircraft sizing problem now becomes not just a root-finding exercise, but also a constrained optimization problem.

The hybrid electric vehicle sizing problem can be sub-divided into three areas: modelling methods/fidelity, energy management and optimization technique. The literature is reviewed to find desirable characteristics and features of each area. Subsequently, a new process for sizing a new hybrid electric aircraft is proposed by synthesizing techniques from model predictive control and detailed conceptual design modelling. Elements from model predictive control and concurrent optimization are combined to formulate a new structure for the optimization of the sizing and energy management of future aircraft.

While the example optimization formulation provided is specific to a hybrid electric concept, the proposed structure is general enough to be adapted to any vehicle concept which contains multiple degrees of control freedom that can be optimized continuously throughout a mission.

The proposed technique is novel in its application of model predictive control to the conceptual design phase.

This paper aims to review national aeronautics and space administration (NASA’s) broad investments in electrified aircraft propulsion (EAP). NASA investments are guided by an assessment of potential market impacts, technical key performance parameters, and technology readiness attained through a combination of studies, enabling fundamental research and flight research.

The impact of EAP varies by market and NASA is considering three markets as follows: national/international, on-demand mobility and short-haul regional air transport. Technical advances in key areas have been made that indicate EAP is a viable technology. Flight research is underway to demonstrate integrated solutions and inform standards and certification processes.

A key finding is that sufficient technical advances in key areas have been made, which indicate EAP is a viable technology for aircraft. Significant progress has been made to reduce EAP adoption barriers and further work is needed to transition the technology to a commercial product and improve the technology, so it is applicable to large transonic aircraft.

Significant progress has been made to reduce EAP adoption barriers and further work is needed to transition the technology to a commercial product and improve the technology, so it is applicable to large transonic aircraft.

This paper will review the activities of the hybrid gas-electric subproject of the Advanced Air Transport Technology Project, the Revolutionary Vertical Lift Technology Project and the X-57 Flight Demonstration Project, and discuss the potential EAP benefits for commercial and military applications. This paper focuses on the vehicle-related activities, however, there are related NASA activities in air space management and vehicle autonomy activities, as well as a breakthrough technology project called the Convergent Aeronautics Solutions Project. The target audience is people interested in EAP.

The purpose of this article is to provide an outline of the challenges of thermal management for more-electric, hybrid-electric and all-electric aircraft, and to notionally discuss potential solutions.

A code algorithm was developed to facilitate architecture-level analysis of the coupled relationship between the propulsion system, the thermal management system, and the takeoff gross weight of aircraft with advanced propulsion systems.

A variety of coupled relationships between the propulsion and thermal management systems are identified, and their impact on the conceptual design choices for electric aircraft are discussed qualitatively.

This conceptual article merely illuminates some driving factors associated with thermal management. The software is still in its adolescence and is experiencing ongoing development.

Thermal regulation in electric aircraft is shown to be a topic that should be addressed in tandem with propulsion system architecture definition and component selection. High-power electronics are expected to emit an immense amount of heat, and the common avenues of heat dissipation could substantially impact the aircraft’s weight, drag and performance. Conversely, strategic management of this waste heat could support subsystems or even produce additional thrust.

This paper aims to direct the attention of researchers and designers in the field of hybrid- or all-electric aircraft design toward the challenges and potential benefits of thermal management.

This paper describes a novel conceptual design software and discusses its logic flow and implications.