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This paper aims to create a terminal area operations (TAO) analysis software that can accurately appreciate the nuances of hybrid electric distributed propulsion (HEDP), including unique failure modes and powered-lift effects.

The program was written in Visual Basic with a user interface in Microsoft Excel. It integrates newly defined force components over time using a fourth order Runge-Kutta scheme.

Powered-lift, HEDP failure modes and electrical component thermal limitations play significant roles on the performance of aircraft during TAO. Thoughtful design may yield better efficiency; however, care must be given to address negative implications. Reliability and performance can be improved during component failure scenarios.

This program has and will support the investigation of novel propulsion system architectures and aero-propulsive relationships through accurate TAO performance prediction.

Powered-lift and HEDP architectures can be employed to improve takeoff and climb performance, both during nominal and component failure scenarios, however, reliance on powered-lift may result in faster approach speeds. High-lift and system failure behavior may also allow new approaches to design and sizing requirements.

This program is unique in both the public and private sectors in its broad capabilities for TAO analysis of aircraft with HEDP systems and powered-lift.

The purpose of this paper is to explore some of the challenges associated with the integration of an LH2-fuelled advanced hybrid-electric distributed propulsion system with the airframe. The airframe chosen as a case study is an ultra-high-capacity blended wing body configuration. It is designed to represent an A-380 class vehicle but in the 2025-2030 timeframe. The distributed propulsion system is a hybrid-electric concept that utilizes high-temperature superconducting technologies. The focus of the study is the application of LH2 as a fuel, with comment being given to kerosene and LCH4.

The study consists of a conceptual design developed through the preliminary design phase and part way into the detailed design phase.

The relationship between passenger capacity and fuel capacity is developed. Some remaining challenges are identified.

The study supports further conceptual design studies and more detailed system studies.

The study contributes to the development of more environmentally benign aviation technologies. The study may assist the development of solutions to the peak oil challenge.

The study explores the integration of a number of complex systems into an advanced airframe to an unusual depth of engineering detail.

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 research is to explore innovative aircraft concepts that use flexible wings and distributed propulsion to significantly reduce fuel burn of future transport aircraft by exploiting multidisciplinary interactions.

Multidisciplinary analysis and trajectory optimization are used to evaluate the mission performance benefits of flexible wing distributed propulsion aircraft concepts.

The flexible wing distributed propulsion aircraft concept was shown to achieve a 4 per cent improvement in L/D over a mission profile consisting of a minimum fuel climb, minimum fuel cruise and continuous descent.

The technologies being investigated may lead to mission adaptive aircraft that can minimize drag, and thus fuel burn, throughout the flight envelope.

The aircraft concepts being explored seek to create synergistic interactions between disciplines for reducing fuel burn while capitalizing on the potential benefits of lightweight, flexible wing structures and distributed propulsion.

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.

The purpose of this paper is to describe the necessity for the use of fully superconducting electrical power systems (SEPS) in future hybrid electric aircraft which facilitates the use of a distributed propulsion system.

The paper looks at the overall design of the electric power systems for these applications and compares the design process of a more conventional power network with a fully superconducting one. The design issues and solutions in each case are then described.

The paper concludes that SEPS will give many advantages to the aircraft design and operation.

Significant efforts needs to be oriented towards the development of fully SEPS and dedicated facilities are required for reliable experimental data that will allow the modelling of these systems.

The requirement for more experimental work has not yet been considered by the Industry, as it is a general belief that these networks will behave similar to the conventional ones.

Recent advancements in electrified transportation have been necessitated by the need to reduce environmentally harmful emissions. Accordingly, several aviation organisations and governments have introduced stringent emission reduction targets for 2050. One of the most promising technologies proposed for achieving these targets is turboelectric distributed propulsion (TeDP). The objective of this study was to explore and identify key indicators for enhancing the applicability of TeDP in air transportation.

An enhancement valuation method was proposed to overcome the challenges associated with TeDP in terms of technological, economic and environmental impacts. The result indicators (RIs) were determined; the associated performance indicators (PIs) were analysed and the key RIs and PIs for TeDP were identified. Quantitative measurements were acquired from a simulated TeDP case study model to estimate the established key PIs.

It was determined that real-world TeDP efficiency could be enhanced by up to 8% by optimising the identified key PIs.

This study is the first to identify the key PIs of TeDP and to include a techno-economic environmental risk analysis (TERA) based on the identified key PIs. The findings could guide developers and researchers towards potential focus areas to realise the adoption of TeDP.

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.

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.