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The purpose of this paper is to define reliability requirements to be imposed on electric engines to assure similar or higher value of mean time between failures (MTBF) for mixed piston-electric propulsion configurations when compared to classic and unconventional piston engine configurations.

Reliability estimation was done using mathematical model of safety of light aircraft commercial operations. The model was developed on the basis of Federal Aviation Administration and National Transport Safety Board data. The analysis was conducted for numerous piston and electric configurations. It allowed comparison of selected solutions and definition of relation between electric engine MTBF and MTBF calculated for entire mixed piston-electric propulsion system.

It was found that, from reliability point of view, mixed piston-electric engine propulsion is attractive alternative for classic single- and twin-piston configuration. It would allow to at least doubling of MTBF for propulsion without increase of operational cost.

Rationale behind exploiting electric propulsion in aviation is provided. Relation between electric engine reliability and entire propulsion reliability was identified and defined. Minimum requirements concerning MTBF value for electric engine application in aviation was assessed. Conclusions from this study can be used for definition of requirements for new aircraft and by the regulatory authorities.

Originality consists in use of real accident statistics included in mathematical model of safety for assessment of MTBF for various classic and novel piston and piston-electric engine configurations of light aircraft. Output from the study can be exploited by the industry.

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.

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.

The purpose of this paper is the multi-disciplinary conceptual investigation of a propulsive fuselage (PF) aircraft layout allowing for new performance synergies through closely coupled propulsion/airframe integration. The discussed aircraft layout facilitates the ingestion of the fuselage boundary layer and the utilization of wake filling, thus eliminating a significant share of fuselage drag.

Based on consistent book-keeping standards for conventionally installed and highly integrated propulsion systems, key aspects of conceptualisation regarding airframe and propulsion system are presented. As a result of this, a PF aircraft configuration is proposed featuring a fuselage fan power plant in conjunction with two under-wing podded power plants. Parametric models for integrated aircraft and propulsion system sizing and performance analysis are discussed that are suitable for the consistent mapping of the characteristics intrinsic to a PF layout. In an initial benchmarking exercise, the vehicular efficiency potentials of the previously identified PF configuration are evaluated against an advanced conventional reference aircraft.

During benchmarking, it was found that a best and balanced design for the proposed PF aircraft layout yields an increase in vehicular efficiency of approximately 10 per cent compared to the advanced conventional reference aircraft.

The paper gives the reader an idea for the efficiency potentials achievable through a PF aircraft configuration, as well as guidelines for aircraft sizing and integrational aspects. It may serve as a basis for advanced studies in the future.

The conceptual investigation of the PF concept idea, contributes to establishing the initial technical feasibility of this novel approach to synergistic propulsion system integration. The methods presented in this paper allow for the multi-disciplinary conceptual design sizing of a PF aircraft.

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.

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 propulsion system integration of a turboprop aircraft, which has been developed for the basic trainer, was performed. The proper turboprop engine was selected among worldwide existing engines by the specific developed engine selection technique and trade‐off studies such as customer’s request for operational capability (ROC), propulsion system parameters, performance analysis with engine installed effects, future growth potential, integrated logistic support (ILS), maintainability, interfaces with the airframe, etc. The chin type air inlet with the plenum chamber was designed in consideration of the inclined configuration to minimize the propeller swirl effect, the inertial separation bypass device to reduce FOD, and the super‐ellipse and NACA‐1 profile lip to maximize the ram recovery. The air inlet was analyzed by a higher‐order source panel method considering propeller wake. The exhaust duct was designed through internal cross‐section area determination to maximize the cruising power as well as external configuration to maximize the effective power, to minimize the aerodynamic drag and to minimize the cockpit contamination by the exhaust gas. The proper oil cooler for the selected turboprop engine was determined with cooling requirements and the oil cooling inlet duct with NACA configuration was designed. The test of the propulsion system including the installation performance test with the effects of the air inlet, the exhaust duct, the propeller and the nose fuselage configuration was performed in the test cell.

Boundary layer ingestion (BLI) is one of the probable noteworthy features of distributed propulsion configuration (DPC). Because of BLI, strong coupling effects are generated between the aerodynamics and propulsion system of aircraft, leading to the specific lift and drag aerodynamic characteristics. This paper aims to propose a model-based comprehensive analysis method to investigate this unique aerodynamic.

To investigate this unique aerodynamics, a model-based comprehensive analysis method is proposed. This method uses a detailed mathematical model of the distributed propulsion system to provide the essential boundary conditions and guarantee the accuracy of calculation results. Then a synthetic three-dimensional computational model is developed to analyze the effects of BLI on the lift and drag aerodynamic characteristics.

Subsequently, detailed computational analyses are conducted at different flight states, and the regularities under various flight altitudes and velocities are revealed. Computational results demonstrate that BLI can improve the lift to drag ratio evidently and enable a great performance potentiality.

The general analysis method and useful regularities have reference value to DPC aircraft and other similar aircrafts.

This paper proposed a DPS model-based comprehensive analysis method of BLI benefit on aerodynamics for DPC aircraft, and the unique aerodynamics of this new configuration under various flight altitudes and velocities was revealed.

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.

This paper aims to provide the international aeronautical community with details of the development of a new disruptive technology for aircraft propulsion.

This paper describes the results achieved by a small Austrian aeronautical innovations company in developing a cyclogyro propulsion system capable of vertical launch and efficient forward flight. The research team progressed from concept definition and simulation (2004-2006), through experimental validation and concept demonstration (2006), component optimization (2006-2012), full system demonstration (2012-2014) and examination of ability to scale (both larger and smaller) (2015 onwards). This paper provides details of the results of each of these stages.

The research team proved that cyclogyro propulsion can be used for the vertical launch, and that, in forward flight, it has the potential to achieve efficiency beyond the range of conventional fixed wing and rotorcraft.

This research indicates that the efficiency increases with forward speed within the range achieved in standard wind tunnels (up to 35 m/s). This efficiency appears to be caused by a unique chamber effect within the cyclogyro rotor assembly. Future research should be conducted to analyse this chamber effect in greater detail and to test the cyclogyro rotor for speeds beyond 35 m/s.

This work indicates that cyclogyro propulsion could have the potential to provide vertical launch, high speed and highly efficient aircraft that have reduced wing span, no external rotors and exceptional agility. This technology could therefore be feasible for vertical take-off and landing aircraft that can safely form densely packed swarms.

It could be researched as an efficiency increase in forward flight completely different to existing propulsion systems. This could open a way for a more efficient air traffic in future and faster reduction of CO₂ and NOX emission an allow an environment-friendlier air travelling.

This paper provides the details of the first cyclogyro aircraft to have flown and will serve the aeronautical community by stimulating the debate on this new disruptive technology.