Conceptual design of hybrid-electric transport aircraft

Progress in Aerospace Sciences ∎ (∎∎∎∎) ∎∎∎–∎∎∎

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Conceptual design of hybrid-electric transport aircraft C. Pornet n,1, A.T. Isikveren 2 Bauhaus Luftfahrt, Ottobrunn 85521, Germany

art ic l e i nf o

a b s t r a c t

Article history: Received 20 April 2015 Received in revised form 15 September 2015 Accepted 16 September 2015

The European Flightpath 2050 and corresponding Strategic Research and Innovation Agenda (SRIA) as well as the NASA Environmentally Responsible Aviation N þ series have elaborated aggressive emissions and external noise reduction targets according to chronological waypoints. In order to deliver ultra-low or even zero in-flight emissions levels, there exists an increasing amount of international research and development emphasis on electrification of the propulsion and power systems of aircraft. Since the late 1990s, a series of experimental and a host of burgeouning commercial activities for fixed-wing aviation have focused on glider, ultra-light and light-sport airplane, and this is proving to serve as a cornerstone for more ambitious transport aircraft design and integration technical approaches. The introduction of hybrid-electric technology has dramatically expanded the design space and the full-potential of these technologies will be drawn through synergetic, tightly-coupled morphological and systems integration emphasizing propulsion – as exemplified by the potential afforded by distributed propulsion solutions. With the aim of expanding upon the current repository of knowledge associated with hybrid-electric propulsion systems a quad-fan arranged narrow-body transport aircraft equipped with two advanced Geared-Turbofans (GTF) and two Electrical Fans (EF) in an under-wing podded installation is presented in this technical article. The assessment and implications of an increasing Degree-of-Hybridization for Useful Power (HP,USE) on the overall sizing, performance as well as flight technique optimization of fuelbattery hybrid-electric aircraft is addressed herein. The integrated performance of the concept was analyzed in terms of potential block fuel burn reduction and change in vehicular efficiency in comparison to a suitably projected conventional aircraft employing GTF-only propulsion targeting year 2035. Results showed that by increasing HP,USE, significant fuel burn reduction can be achieved; however, this also proves to be detrimental in terms of vehicular efficiency. The potential in block fuel reduction diminishes with increasing design range – especially for low battery gravimetric specific energies. In addition, the narrow shape of the fuselage represents a volumetric constraint for the storage of the battery and typical cargo. It was concluded that the short-range/regional market segment would be the most suited for the application of such concepts. Concerning the influence of HP,USE on flight technique optimization, an increasing HP,USE was found to have a tendency of decreasing the optimum flight speed and altitude. Further investigation of more synergistic design and integration of the hybrid-electric motive power system needs to be conducted in order to explore the full benefit of such technologies. & 2015 Elsevier Ltd. All rights reserved.

Keywords: Hybrid-electric transport aircraft Fuel-battery hybrid Aircraft design Aircraft sizing Integrated performance

Contents 1. 2.

3.

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Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Aircraft morphologies and systems architectures for electro-mobility . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6 2.1. Topological options, taxonometric conventions and algebraic descriptors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6 2.2. Integrated propulsion and power for fixed-wing aviation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7 2.2.1. Experimental and commercial activities . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7 2.2.2. Synergy with distributed propulsion and survey of future concepts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 2.3. Scope of investigative work in this article . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9 Sizing scheme for the hybrid-electric propulsion system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

Corresponding author. 1 Researcher, Integrated Hybrid-Energy Propulsion and Power Systems, Visionary Aircraft Concepts. 2 Head, Visionary Aircraft Concepts.

http://dx.doi.org/10.1016/j.paerosci.2015.09.002 0376-0421/& 2015 Elsevier Ltd. All rights reserved.

Please cite this article as: C. Pornet, A.T. Isikveren, Conceptual design of hybrid-electric transport aircraft, Progress in Aerospace Sciences (2015), http://dx.doi.org/10.1016/j.paerosci.2015.09.002i

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3.1. Flow path sizing of the Turbofan and Electrical Fan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 3.2. Turbofan model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 3.3. Ducted fan model and sizing strategy of the electric motor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 3.4. Discussion about the model and its limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11 3.5. Electrical system characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12 4. Aircraft sizing and integrated performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12 4.1. Parametric descriptors and qualitative pre-design investigation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12 4.2. Aircraft Top-level Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13 4.3. Design mission. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13 4.4. Aircraft sizing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14 4.4.1. Sizing guidelines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14 4.4.2. Electric motor critical sizing conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14 4.5. Relative change in block esar versus relative change in block fuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14 4.6. Relative change in block COSAR versus relative change in block fuel. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15 4.7. Relative change in MTOW versus the Degree-of-Hybridization for Useful Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15 4.8. Relative change in point ESAR versus the Degree-of-Hybridization for Useful Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16 4.9. Influence of Degree-of-Hybridization for Useful Power on the combustion-based propulsion system efficiency . . . . . . . . . . . . . . . . . . . 16 4.10. Influence of the Degree-of-Hybridization for Useful Power on take-off performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17 4.11. Sensitivity study according to battery gravimetric specific energy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17 5. Influence of the Degree-of-Hybridization for Useful Power on flight technique optimization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17 6. Aircraft characteristics and benchmark . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19 6.1. Aircraft benckmark . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19 6.1.1. Aircraft main data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19 6.1.2. Aircraft mass breakdown. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19 6.2. Optimum flight technique analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19 6.2.1. Reference aircraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20 6.2.2. Hybrid-electric aircraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21 7. Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21 Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21 Appendix A. Supplementary material. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21 References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

1. Introduction viation today represents 2% of anthropometric carbon dioxide (CO2) emissions [1]. Objectives for Vision 2020 of the Advisory Council for Aeronautics Research in Europe (ACARE) target an 80% and 50% reduction in nitrous oxide (NOx) and CO2 respectively [2]. Even more ambitious goals outlined in Flightpath 2050 [3] by the European Commission (EC) for year 2050 is a 75% reduction in CO2-emissions per passenger kilometer (PAX.km) relative to the capabilities of conventional aircraft of the year 2000. Furthermore, a 90% reduction of NOx-emissions and a 65% perceived noise reduction is advocated. Finally, aircraft movements on the ground have to be emission-free when taxiing. The scope of the Flightpath 2050 assessment comprises total emissions between leaving the parking position at an origin airport (off-block) and the arrival at position at the final destination (on-block). From an international perspective one can compare and contrast these EC objectives to those espoused by the International Air Transport Association [4] by way of the Air Transport Action Group [5], the International Civil Aviation Organization [6] and the US National Aeronautics and Space Administration [7]. Irrespective of the agenda or governmental office in question the conclusion is that all these targets call for a dramatic reduction in emissions over the interim-to-long term. Targets for CO2-emissions as originally defined in Vision 2020 and AGAPE 2020 [8] were categorized into Airframe, Propulsion and Power System (PPS), Air Traffic Management (ATM) and Airline Operations. As exemplified by Fig. 1, the Strategic Research and Innovation Agenda (SRIA) goals [9] have been re-calibrated to reflect the achievements assessed by the AGAPE 2020 report and a new medium-term goal for Year Entry-into-Service (YEIS) 2035, which is a significant point for aircraft fleet renewal. A further elaboration of the chronologically assigned CO2-emissions targets

is a breakdown that recommends aircraft energy levels (for flight including all on-board systems and services). As shown in Fig. 2, the NASA Environmentally Responsible Aviation N þ series targets [7] apply to technology freeze year as opposed to YEIS espoused by SRIA and Flightpath 2050. A technology freeze year infers attainment of Technology Readiness Level [10] (TRL) 6, i.e. primed for a product development programme, and generally, an interval of at least 5 years would characterize technology freeze and YEIS milestones. If one peruses the various targets set by N þ3, the stated fuel/energy consumption (proxy for CO2-emissions) reduction of 60% is synonymous with the goal set by SRIA 2035. This means the N þ3 target can be considered to be somewhat aggressive compared to the European goal in a temporal sense. A similar conclusion can be drawn when conducting a comparison of Landing–Takeoff cycle (LTO) NOx-emissions targets. In keeping with the review conducted by Isikveren and Schmidt [11], focusing on the SRIA goals, in order to realize a total 60% reduction in fuel burn and corresponding CO2-emissions per PAX.km for target YEIS 2035, SRIA 2035 [9] suggests 51% from combined Airframe and PPS, and, 9% from improved ATM and operational efficiency. If one extends beyond year 2035 in order to consider a plausible strategy for Flightpath 2050, according to SRIA 2050 a possible breakdown for the total 75% reduction in CO2-emissions would be 68% from Airframe and PPS combined, and, the remaining 7% of this 75% total from improvements through ATM and Operations. Data compiled from various investigations have shown the PPS category demands an efficiency improvement of approximately 80% over the reference year 2000 [12]. Advanced gas-turbine concepts, such as those based upon the classic Joule-Brayton cycle through intercooling and recuperation is expected to lead to thermal efficiencies of around 50% or even slightly higher, and, a further overall gain might be realized

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Nomenclature Symbols a local sonic velocity (m/s) COSAR Cost Specific Air Range (nm/USD) CDo zero-lift drag coefficient (dimensionless) dCD/dCL2 vortex-induced drag factor (dimensionless) e battery gravimetric specific energy at cell-level (Wh/ kg); Euler's number E energy (Wh) ESAR energy Specific Air Range (nm/kWh) g acceleration due to gravity (m/s2) HE Degree-of-Hybridization for energy (dimensionless) HP Degree-of-Hybridization for power (dimensionless) L/D lift-to-drag ratio (dimensionless) M Mach number (dimensionless) N number of propulsors (dimensionless), corrected fan speed (dimensionless) P power (W) SW reference wing area (m2) T thrust (N); time (min) TSFC Thrust Specific Fuel Consumption (g/kNs) TSPC Thrust Specific Power Consumption (W/N) W instantaneous gross weight (kg) η overall power plant efficiency (dimensionless) ηM rate change of overall power plant efficiency w.r.t. Mach number (dimensionless) s local density lapse ratio (dimensionless) ϕ Activation ratio (dimensionless) Φ Supplied Power Ratio (dimensionless) Subscripts EF ELEC INS LRC MRC rel TF TOC TOT USE

Electric fan electrical installed [Power] Long Range Cruise Maximum Range Cruise relative Turbofan Top-OF-Climb Total Useful [Power]

Acronyms and abbreviations ACARE AEO ATAG ATLeRs

Advisory Council for Aviation Research and Innovation in Europe All-Engines Operational Air Transport Action Group Aircraft Top-Level Requirements

through ultra-high by-pass ratio ducted fans and the Open Rotor. Other published advanced studies indicate the reduction due to Airframe will not offer more than around 25% [13–15] of this 68% total. Even factoring in an aggressive development strategy for combustion based PPS something like a 10–15% CO2-emissions is still left unaccounted for. Tellingly, it can be concluded that electrification of PPS could have the potential to deliver the target for year 2035. Beyond this, in order to achieve ultra-low or even zero in-flight emissions levels of energy hybridity tending towards a much higher proportion of electrification appear necessary.

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ATM Air Traffic Management BCU Battery Controller Unit BHL Bauhaus Luftfahrt e.V. BLI Boundary Layer Ingestion CO2 carbon dioxide CU Cranfield University DEAP Distributed Electrical Aerospace Propulsion DESPPS Dual-Energy Storage-Propulsion-Power System DF Ducted-Fan DisPURSAL Distributed Propulsion and Ultra-high By-pass Rotor Study at Aircraft Level DoH Degree-of-Hybridization DWB Discrete Wing and Body (“tube-and-wing”) EC European Commission EF Electric Fan EM Electric Motor FL flight Level, hundreds of feet pressure altitude GTF Geared-Turbofan GW Gross Weight HTS High Temperature Superconducting ICAO International Civil Aviation Organization IGW Increased Gross Weight IATA International Air Transport Association ISA International Standard Atmosphere LERC Long ESAR Range Cruise LRC Long Range Cruise LTO Landing-Takeoff cycle MCRC Maximum COSAR Range Cruise MERC Maximum ESAR Range Cruise MTOW Maximum Take-Off Weight MLW Maximum Landing Weight MRC Maximum Range Cruise NASA National Aeronautics and Space Administration NOx nitrous oxide OEI One Engine Inoperative OEW Operating Empty Weight PAX Passengers PMAD Power Management and Distribution PPS Propulsion and Power System SAR Specific Air Range SL Sea Level SRIA Strategic Research and Innovation Agenda SSPC Solid State Power Controller TeDP Turbo-electric Distributed Propulsion TF Turbofan TOFL Take-Off Field Length TOC Top-Of-Climb TRL Technology Readiness Level USD US Dollars USG US Gallons YEIS Year Entry-into-Service

Fundamentally, ideas first proposed by Sir George Cayley [16] in the turn of the 19th Century have been the principles governing aerospace vehicle design up to this point in time. The synthesis of a complex mechanical, airborne transportation system typified by multiple interactions and often conflicting requirements were, and still are, successfully approached by breaking the system up into disparate, weakly coupled entities. In an effort to promote further improvement in product development outcomes, contemporary design teams in industry comprise individuals who serve in multidisciplinary sub-groups; for instance amongst others, wing-

Please cite this article as: C. Pornet, A.T. Isikveren, Conceptual design of hybrid-electric transport aircraft, Progress in Aerospace Sciences (2015), http://dx.doi.org/10.1016/j.paerosci.2015.09.002i

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C. Pornet, A.T. Isikveren / Progress in Aerospace Sciences ∎ (∎∎∎∎) ∎∎∎–∎∎∎

Fig. 1. Chronologically defined CO2 and NOx-emissions reduction goals as recommended by ACARE in the SRIA document [9].

fuselage, airframe-engine and integrated utilities. Although proven beneficial, this approach is hampering any significant potential for further optimization because it is curtailing what is now considered to be a restrictive design space. Advanced concepts require treatment of the design problem in a holistic sense, with emphasis placed upon maximizing synergy in the global system. Synergy is conception of a particular system that serves to benefit or ameliorate other vehicular systems in a cross-disciplinary manner. In the context of transport aircraft design, it was argued earlier that future aircraft product development will rely on significant strides in PPS research – and there exists growing evidence hybrid-electric, or even so-called universally-electric (fully-electric propulsion with All Electric Aircraft major systems) solutions are the logical options. This means any research activity addressing questions related to emissions and noise reduction for aviation should to some extent exhibit synergy or complement initiatives

undertaken for PPS. In order to ensure the continued relevance and success of all such research activities the realization of ultralow emissions systems infers it should be treated as a tightlycoupled multi-faceted problem, with the implication any activities dealing with new architectural approaches should cover: 1. [Hybrid-] Electrical energy generation and/or storage, including synergy with airframe; 2. Fully Integrated Power Management and Control (including actuation and Flight Control System); and, 3. Integrated Utilities, i.e. major-systems like Environmental Control System, Landing Gear and Avionics. As a further requirement, due to the inter-disciplinary nature of such investigative work understanding and evaluating localized systems attributes at aircraft level becomes essential. Another aspect that is of prime importance is a set of tenets that govern the acceptance and longevity of new technologies

Fig. 2. Chronologically defined CO2 and NOx-emissions, as well as community noise reduction targets as recommended by NASA in the Environmentally Responsible Aviation document [7].

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utilized for transport aircraft. Aside from complexities related to certification and overcoming the inertia associated with adopting new practices by suppliers and airworthiness authorities alike, it can be argued that any technology for YEIS 2025 þ requires four prime considerations:

 Although it may not deliver the potential peak attributes of performance by target YEIS, it still can provide a solid business

 

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case at any point in time, e.g. as in the case for More-Electric or All-Electric Aircraft components, sub-systems and architectures; Possess attributes where scope for performance improvement exists over an intermediate-to-long-term period even after initial service-entry, i.e. allows for evolutionary development; Is scalable whilst still retaining performance such that in-house knowledge can be migrated from one product development programme to another with relative ease; and,

Fig. 3. (A) Conventional, hybrid-electric and universally-electric propulsion system architectural options, source [17]; (B) Possible power-train options for hybrid-electric propulsion, upper portion denotes serial arrangements, lower portion denotes parallel arrangements, key: G – generator/M – motor.

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Fig. 4. Example of a Degree-of-Hybridization trade-study conducted for a hypothetical Dual-Energy Storage-Propulsion-Power System targeting YEIS 2035; source Isikveren et al. [19].

 Maximizes synergy such that as many aircraft systems as possible are improved.

2. Aircraft morphologies and systems architectures for electro-mobility The aviation sector can be considered to be in the midst of a pioneering era with regards to electro-mobility. As a result, dramatic and disruptive changes to component/sub-systems technologies coupled with experimentation of PPS architectures and/ or aircraft morphologies is currently taking place. The purpose of this section is to lend visibility to what options are available to hybrid-electric and universally-electric aircraft solutions, to present state-of-art commercial activities, and to review the current array of future concepts proposed by academia and industry. 2.1. Topological options, taxonometric conventions and algebraic descriptors A comprehensive representation of the most relevant combinatorial variety of hybrid-electric propulsion system for aircraft transport application can be proposed by distinguishing between the components (or component chains) generating the shaft power and the devices consuming it, as illustrated in Fig. 3A [17]. The term hybrid-electric implies intrinsically that electric power is used in combination to a least one additional power source (commonly fuel power). Assuming regarding the power-train a combustion engine and an electrical source, the first step is to establish whether they are combined in serial or in parallel. This distinction can be tied to the nature of the power node between the system constituents: in a serial hybrid arrangement, the node is electrical (upper portion of Fig. 3B), while in a parallel hybrid, it is mechanical (lower portion of Fig. 3B).

The Degree-of-Hybridization (DoH) employed in such advanced systems cannot be suitably represented by a single parametric descriptor. Lorenz et al. [18] have argued a full description of any generic hybrid-PPS requires two descriptors involving account of both the alternative energy [source] and that of the entire PPS: one ratio comparing each of the maximum installed (or useful) powers (HP); and, a second ratio comparing the extent of energy storage (HE) of each, viz.

HP =

P ELEC E and HE = ELEC P TOT ETOT

For a hybrid-electric solution, PELEC would represent the maximum installed (or useful) electric power, and PTOT the total PPS installed power (motor, and for example, gas-turbine), EELEC the total stored electric energy, and ETOT the total stored energy for the entire PPS (electrical, and for example, kerosene). In order to elucidate why such a dual set of parametric descriptors are necessary, consider:

 Conventional kerosene based gas turbine PPS – here HP ¼0 and HE ¼0; or,

 Pure serial hybrid-electric architecture where only electrical 

power is provided at the propulsive device(s) but energy storage is solely kerosene based – here HP ¼1 and HE ¼ 0; or, Universally-electric aircraft where energy storage is batteries only – here HP ¼1 and HE ¼ 1.

Isikveren et al. [19] established an algebraic basis when describing the DoH for Power, HP, and, DoH for Energy, HE, parametric descriptors for purposes of interpreting the results of advanced trade-studies, optimization and corresponding synthesis of coherent engineering solutions involving any type of Dual-Energy Storage–Propulsion–Power System (DESPPS). They reasoned for a given set of standalone sub-system energy conversion efficiencies,

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Fig. 5. A survey of experimental, in-production or near in-production universally-electric aircraft sequenced according to year of first flight; first manned, fixed-wing electric flight shown at apex of chart; information compiled from [28,29].

the parametric descriptor of HP was found to be solely a function of the Supplied Power Ratio (Φ, related to converted power afforded by each energy carrier), whereas in contrast, HE was found to be a more complex synthetic function described by comingling of Φ and the Activation Ratio (ϕ, describing the relative nature of utilization with respect to time afforded by the motive power device associated with each energy source). The non-dimensional parameters Φ and ϕ each vary between zero (denoting utilization of Energy Source “a” only) and unity (denoting utilization of Energy Source “b” only). For a given DESPPS architecture and inherent attributes related to component and sub-systems efficiencies this analytical representation can be visualized using the format of a Ragone diagram [20]. Fig. 4 above displays an example functional correlation between installed HP and HE for a DESPPS based upon kerosene and batteries as energy carriers targeting YEIS 2035 and were derived from assumed step values of Φ and ϕ. It has been shown when sizing the electrical part of a hybrid-electric system, the batteries for instance, that the energy or the power requirement can represent the sizing criterion [21]. Since for electrochemical storage devices like batteries, gravimetric specific power and gravimetric specific energy are not independent, this means that the power sizing simultaneously influences the amount of electrical energy stored. In contrast, energy and power sizing are independent for fuel based systems, because the gravimetric specific power of fuel is unlimited, while the energy amount stored is directly proportional to the fuel mass. Another aspect of the work conducted by Isikveren et al. [19] was to establish which aircraft morphologies would appropriately match a hybrid-electric PPS approach. One of the important outcomes of any pre-design exercise is to establish whether a given advanced systems approach would be further enhanced by the adoption of aircraft morphologies other than the traditional so-called “tube-andwing” combination. An opportunity to gauge attributes related to vehicular aerodynamic efficiency and its association with DESPPS was afforded when potential useful loads between the aircraft body and wing were judiciously shifted and from a high-speed aerodynamics perspective the relative merits between what is

termed the “Discrete Wing and Body” (DWB, also referred to as the “tube-and-wing”), the “Integrated Wing Body” and the “All-Wing Aircraft” were inspected. By comparing the ratio of wing displaced volume to the total displaced volume of the aircraft, this independent parametric variable posed opportunity to gauge whether all solutions generally fall within the bounds of what was defined as a DWB, or, a “tube-and-wing” aircraft morphology. Results found for working parameters typical of aircraft employing hybrid-electric PPS undertaking short-haul operations, the sizing outcome of the relationship between wing displaced volume to that of the total displaced volume of aircraft was more indicative of a contemporary ultra-large, ultra-long haul aircraft. This meant, unless a significant departure in PPS integration is considered, e.g. distributed propulsion, or, a universally-electric PPS is selected from the outset, the tube-and-wing morphology was still considered to be appropriate. 2.2. Integrated propulsion and power for fixed-wing aviation Progress from the first ever manned, fixed-wing flight of an aircraft solely using electrical energy for the PPS has been relatively slow in ramping up until around the year 2000 [22–27]. A number of commercial offerings exist; however, major inroads still need to be made for such technology to be incorporated in dedicated Part 25 transport category aircraft. Here, a number of commercial activities will be briefly reviewed and synergistic design morphologies, like distributed propulsion, will be discussed. This section motivates why a hybrid-electric, “tube-and-wing” aircraft morphology with quad-fan arrangement would complement the database of future concepts being studied today. 2.2.1. Experimental and commercial activities At this point in time, electro-mobility for aircraft only exists in the single/twin-seater categories and are typically retrofits of existing conventional designs with reduced payload-range capability. Their construction is often motivated by technical curiosity; however, there exists an emerging commercial interest limited to

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glider, ultra-light and light-sport airplane markets. The era of socalled “electric aircraft” (or perhaps more appropriately “universally-electric” as suggested by the authors of this article earlier) had its beginnings with Brditschka’s converted HB-3 dubbed the MB-E1 [30]. The aircraft, which flew for just over 9 min in 1973, was powered by a 10 kW Bosch motor and utilized Ni–Cd battery cells for energy. The 1970s until around the end of the 1990s were witness to several experimental aircraft, but most of these focused on solar as a source of electric energy. From around the early 2000s until today, a proliferation of experimental aircraft utilizing electro-chemical means of energy supply has emerged. It is around the same time a series of glider/ultra-light/light-sport aircraft integrators, namely, as itemized from left-to-right in Fig. 5, Air Energy, Electraflyer, Yuneec International, Sonex Aircraft, SchemppHirth Flugzeugbau, Lange Aviation, Electravia, PC-Aero, Pipistrel and Cessna have undertaken the step of offering or are scheduled to offer production aircraft to the light aviation sector. Fig. 5 also presents images for each of these production aircraft and has been sequenced according to year of first flight. Apart from a novel butterfly (or Vee-tail) empennage arrangement for the Yuneec E430 and Sonex E-Flight Waiex, all aircraft appear to employ a conventional morphology, indicative of contemporary glider/ultralight/light-sport offerings. Regarding fixed-wing commercial aviation, at current technology levels the development of even a hybrid-electric passenger aircraft appears challenging. Even setting aside the qualitatively different power requirements for low-speed operations, limitations in range and flexibility in the payload-range working capacity (useful load trade) is one of the major impediments. This is exemplified by basic studies conducted on the Dornier Do328 regional aircraft modified into a universally-electric aircraft where the outcome indicated with improved aerodynamics, reduced structural mass still more advanced batteries are necessary to achieve even a modest range capability [31]. Irrespective of these challenges, one of the world's leading commercial aircraft integrators, Airbus Group, has recognized the potential of electromobility for aeronautical applications, and has committed resources to the so-called “E-Aircraft Programme” [29]. It is an industry lead initiative conceived by Airbus Group in order to pave the way forward in a piece-meal fashion for commercial aircraft to operate with fully-electric propulsion as the ultimate goal. At this moment in time, active projects include: the Dimona DA36 e-Star 2nd Generation Project, a two-seat hybrid-electric aircraft in

conjunction with Diamond Aircraft and Siemens AG (see Fig. 5), the E-Fan (see Fig. 5), a fully electrically-powered pilot training aircraft and the E-Thrust concept study based on a hybrid[universally]-electric distributed propulsion system architecture (see Fig. 6). 2.2.2. Synergy with distributed propulsion and survey of future concepts The introduction of electric and hybrid-electric technology has drastically opened the design space. The full-potential of these technologies will be drawn through synergetic integrations of the propulsive system at aircraft level which may lead to rethink entirely the aircraft design paradigm [32]. As conceptual designs for distributed propulsion developed throughout the years, e.g. Distributed Fans and so-called Propulsive Fuselage concepts undertaken in the recently completed EC Framework Programme 7 DisPURSAL Project [33], they have evolved into concepts where the power generated through core gas turbine engines is transmitted to distributed propulsor fans which produced aircraft thrust. The challenge, however, was to establish the most efficient method of transmitting power from the core engine to the propulsor fans. Of the several concepts investigated, the best alternatives included either employing mechanical transmission or producing electrical power and further transmitting it via a conventional, i.e. non-High-Temperature Super-conducting (HTS) technologies, distribution network to several electrically driven fans using electrical motors. It was, however, found that if conventional methods of power transmission were to be utilized, they would result in low efficiency factors, thereby, leading to poor performance and hence higher fuel consumption than conventional systems [23,34]. The key benefits of this synergistic application include achieving ultra-high effective bypass ratios, while retaining the superior efficiency of large core engines, enabling Boundary Layer Ingestion (BLI) to achieve higher propulsive efficiencies, lower fuel burn, reduced level of low-speed operations noise and overall environmental emissions, various integration advantages, improvement in transient operating characteristics of the integrated system and improved benefits for large passenger aircraft over long-range missions. The benefits of the concept additionally allow new degrees of design freedom. These include the location of large turbo-shaft engines driving generators for optimal performance, embedding of a broad continuous array of motor-driven fans on

Fig. 6. A survey of hybrid-electric and universally-electric aircraft concepts targeting YEIS 2030 þ unveiled in recent years, sorted left-to-right according to increasing passenger accommodation.

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the upper surface of the aircraft to maximize propulsive efficiency by ingesting thick airframe boundary layer flow and desired/optimal combination of thrust split between the core propulsion units and boundary layer ingesting fans. Distributed propulsion concepts being pursued today are based on the synergistic application of hybrid-electric architectures and distributed propulsion in a purely serial architectural arrangement – termed as Turbo-electric Distributed Propulsion (TeDP). TeDP as a concept considers the application of super-conductivity and HTS technologies to reduce electrical power transmission losses. The concept has gained momentum in recent years with extensive research currently being undertaken in Europe and the US. The NASA N3-X [35] and the Cranfield University (CU) BW-11 [34] concept aircraft are typical examples of this approach. During the last six years, CU has engaged in a number of projects concerning distributed propulsion. The work has resulted in a number of interesting design studies [36–39] aimed at improving the technology and its development for future implementation. In another effort to distribute thrust, Airbus Group Innovations and RollsRoyce together with CU as are currently engaged in the Distributed Electrical Aerospace Propulsion (DEAP) Project [40]. The E-Thrust conceptual work discussed earlier constitutes one aspect of this broader DEAP initiative. Fig. 6 presents various hybrid-electric and universally-electric aircraft concepts unveiled in recent years sorted according to size (left-to-right), ranging from 4 seats (top-left) up to 555 passengers (bottom-right). All aircraft target a YEIS of 2030 þ. Hybrid-electric concepts include the NXG-50 [41] (50-passenger), DEAP [40] (around 100-passenger), ESAero TeDP [42] (150-passenger), SUGAR Volt [43] (154-passenger), Bauhaus Luftfahrt (BHL) Twin-Fan [44] (180-passenger), BHL Quad-Fan (180-passenger, the focus of this technical article), N3-X [35] (300-passenger) and BW-11 [34] (555-passenger). With the exception of the ESAero TeDP, N3-X and BW-11, all the listed hybrid-electric concepts employ some form of serial/parallel combinatorial power-train arrangement that utilize either a gas-turbine core with generators or advanced batteries as the main source of electrical energy. The ESAero TeDP, N3-X and BW-11 employ a purely serial power-train, i.e. HP ¼ 1 and HE ¼0, consisting of a turbo-shaft engine which is used solely to provide electrical power through a generator to electric motors driving multiple propulsive fans, which are distributed above, below, or inside a wing. The universally-electric concepts, namely, HP ¼1 and HE ¼ 1, are cited as LEAPTech [45] (4 passengers, pilot included), 328-LBM [31] (around 30-passenger), VoltAir [46] (regional, unspecified accommodation capacity) and the Ce-Liner [47] (189passenger). Each of these concepts utilize electrical energy supplied from advanced batteries combined with propulsors either reflecting suitably projected conventional electric motors or motors incorporating HTS technologies. The universally-electric technology is married with distributed technology in the LEAPTech concept [45]. By distributing the thrust on 18 propellers mounting on the wing leading-edge, the design of the wing could be performed to achieve most efficient cruise while still providing similar low-speed performance of a conventional gas-powered aicraft. This concept exemplifies possible synergies which can be drawn through integration of electric propulsion at aircraft level and the change in paradigm for aircraft design. As one reviews the aircraft concepts shown in Fig. 6 above, certain clusters of design intent and integration strategies could be identified. The LeapTech and ESAero TeDP all employ distributed propulsion morphological approaches that not only focus on zero and ultra-low-emissions respectively, but also provide Short Takeoff and Landing (STOL) capabilities. The DEAP and VoltAir exploit benefits afforded by BLI, with the VoltAir producing a net favorable vehicular efficiency outcome through the additional gains by way of Wake Filling, as exemplified by the Propulsive Fuselage Concept

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of the EC Framework Programme 7 DisPURSAL Project [48]. Also, it is interesting to observe apart from the N3-X and BW-11, which have been designed to accommodate a few to several hundred passengers, all concepts assume an advanced “tube-and-wing” arrangement. The N3-X and BW-11 attempt to exploit benefits associated with BLI and noise shielding, and thus the Blended Wing-Body morphology lends itself well to such an approach. 2.3. Scope of investigative work in this article The nature of this technical article is to further examine the relative merits of the advanced “tube-and-wing” aircraft morphology with a hybrid-electric systems approach to the PPS subject to constraints HP≠0,1 and HE≠0,1. It is an extension of previous studies conducted on a hypothetical hybrid-electric narrow-body twin-sized transport targeting YEIS 2035 [21,44] with the main distinction of this work being adoption of a quad-fan arrangement, thus allowing for a feasibility study involving a rudimentary form of distributed propulsion. The complete electrical system integration approach undertaken in previous publications [21,43,44] for the conceptual design of fuel-battery hybrid-electric transport aircraft has been a parallel installation of an Electric Motor (EM) mounted on the shaft of a gas-turbine (GT). The EM can either support the operations of the GT [21,43] or even drive the shaft of the propulsor by itself during some segments of the complete flight mission [44]. Utilizing an EM concurrently with a GT in a parallel installation results in operating the GT in part load, which impairs its efficiency. Driving the propulsor by the EM itself during cruise while the GT is shut down enables one to introduce electric power into the propulsion system while the GT core remains unaffected [44]. However, in this case the DoH for Power (HP) is fixed and cannot be varied. With the aim of assessing the implications of increasing HP on sizing and performance of hybrid-electric aircraft and in the search for an innovative integration of the hybrid-electric motive power system, a new approach is proposed and subsequently investigated in this technical article. Instead of coupling the EM directly to the lowpressure shaft of the GT as in previous work, the EM here is mounted directly on the shaft of the propulsor and this new EMPropulsor entity, dubbed an Electric-Fan (EF), these are integrated in the aircraft as additional bill-of-materials item to the combustion based engines. In order to clarify this approach, one might think of an aircraft with a number of electrical fans, NEF, and a number of Turbofans (TF), NTF. Concrete examples would be a trifan aircraft with two fans conventionally powered by GTs while the remaining fan is driven by an EM, or, as a quad-fan aircraft equipped with two TFs and two EFs. This latter design option is the focus of the investigation presented this paper. As the operations of the EFs and TF are now considered to be independent, the HP can be freely varied and its effect on aircraft design can be assessed. Moreover, an expected advantage of this integration approach is if the sizing of the EFs and TFs are done thoughtfully, the overall power plant efficiency of the TFs will not be significantly affected by the operations of the EFs. The sizing strategy of the hybrid-electric propulsion system with an emphasis on the sizing of the EF is discussed first. The design and performance of a quad-fan concept is then investigated with respect to the enounced sizing scenario of the hybrid-electric propulsion system. The consideration of the overall sizing effects, and in particular, of the integration impact of the electrical system at aircraft level is based on sizing and performance methods previously published by the authors [21]. The implication of increasing HP is discussed for an interval of design ranges. The influence of HP on flight technique optimization is also analyzed. Finally, the “bestand-balanced” hybrid-electric concept is benchmarked against an advanced gas-turbine only aircraft. As a matter of clarification, in

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adopting a convention that permitted an appropriate comparison between terrestrial, marine and aeronautical vehicles [18], and Isikveren et al. [19] presented advanced aircraft trade-study work using the parametric descriptor of HP for installed power (HP,INS). The reader should note for sake of lending transparency and to assist in providing some measure of understanding, all values related to DoH for Power in this technical article adheres to Useful Power as a basis (HP,USE), i.e. the power seen at the propulsive device(s).

3. Sizing scheme for the hybrid-electric propulsion system The sizing scheme for the hybrid-electric motive power system is developed in this section. Flow path sizing of the TFs and the EFs has been performed at Top-Of-Climb (TOC) conditions. The thrust requirement at TOC and the thrust provided by the EFs determine the design thrust of the TFs. Models for the TFs implemented in the quad-fan aircraft overall model is also described. The thrust and power demand characteristics of the Ducted-Fan (DF) are generically introduced to discuss the sizing cases of the EM according to HP,USE. 3.1. Flow path sizing of the Turbofan and Electrical Fan The flow path sizing of the TFs and EFs is performed at TOC condition (ISA, FL350, M0.78). The thrust requirement at TOC, TTOC, which includes a 300 fpm residual climb rate, is distributed between the TFs and the EFs according to HP,USE, the number of EFs, NEF, and the number of TFs, NTF.

TEFTOC =

TTOC⋅Hpuse NEF

TTOC = TTF,TOC⋅NTF + TEF,TOC⋅NEF

A value of zero for HP,USE would represent the case where no EF (s) are sized/utilized in the concept, whereas the case with a value of one would mean a concept where TFs are not sized/utilized. Both of these values are theoretical in the sense that they do not represent a relevant practical engineering solution in the context of the investigated quad-fan approach. 3.2. Turbofan model The TF model is based on a Geared-Turbofan (GTF) model published by Pornet et al. [44] modeled in GasTurb11s [49], it was sized for several design thrusts ranging 15–35 kN at TOC conditions (ISA, FL350, M0.78). For each selected design thrust, geometrical dimensions, weight, maximum thrust and fuel flow characteristics were provided in form of multi-dimensional tables enabling an integrated evaluation of the GTFs at aircraft level. The GTF was designed for an overall pressure ratio of 62.0 and a bypass ratio of 16.2. A gear ratio of 3.0 was found to provide a good compromise between turbine strain and stage loading. The design yielded a Thrust Specific Fuel Consumption (TSFC) of 13.24 g/kNs for the engine with the highest thrust and 13.30 g/kNs for the engine with the lowest thrust. The weight of the engines was calculated according to geometric component dimensions and performance parameters. The TSFC characteristic is represented against the thrust setting for the different engine models. It can be observed that the scale effect has only a small influence on the TSFC characteristics and efficiency. As shown in Fig. 7 an important note for later analysis of the efficiency of the GTFs according to its part load characteristics is that the TSFC bucket covers an interval of 70–100% of the maximum thrust. 3.3. Ducted fan model and sizing strategy of the electric motor

(1) (2)

The required fan diameter of the TFs and EFs is obtained from the design thrust, TTF,TOC and TEF,TOC, according to the TF model and the DF model respectively described in Section 3.2 and in Section 3.3.

The thrust and power-demand characteristics of the EF model are based on a DF model published by Steiner et al. [50] The sizing of the EF which includes the sizing of the DF and of the EM is essential in the design of the quad-fan concept. Consequently, generic thrust and power characteristics of the DF are reviewed in this section in order to explain aspects involved in the sizing. As indicated previously the flow path sizing of the DF is performed at

Fig. 7. Thrust Specific Fuel Consumption (TSFC) characteristics of the Geared Turbofan model with respect to the design thrust at ISAþ 10 °C, FL350 and M0.76.

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16000

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Fig. 8. Ducted-Fan model sized at ISA, FL350 and M0.78 for a design thrust of 10 kN and a fan design pressure ratio of 1.413; (left) off-design power demand at a relative corrected speed of 1.0; (right) off-design thrust characteristics at M0.78.

TOC conditions (ISA, FL350, M0.78). The thrust and power characteristics of a generic DF sized at a design thrust of 10 kN and at a fan design pressure ratio of 1.413 are represented in Fig. 8. The sizing point is indicated by a solid circle. As per definition at the sizing point the relative corrected fan speed, Nrel,EF, is 1.0. The power demand characteristics of the DF are of particular importance for the sizing of the EM. The variation of the power demand with Mach number and altitude are represented in Fig. 8 (left) at Nrel,EF of 1.0. It can be observed first that the off-design power demand increases with decreasing altitude. Secondly, while at high-altitude, the off-design power remains about constant with Mach number, it increases in contrast noticeably with increasing Mach number at low altitude. According to the observation made about the power-demand characteristics of the DF, the sizing scheme of the EM can now be discussed. If the EM is sized according to the DF power required at its sizing point and by recalling the increasing power demand of the DF with decreasing altitude, it becomes clear that the EM installed power will limit the fan performance at lower altitude (nota bene: it is assumed in the model that the EM is thermally managed and that in the first instance its maximal power remains constant with altitude and speed). In other words, the fan speed will be reduced during climb and take-off limiting the available thrust. The decrease of the installed thrust with reduction of the fan speed is represented in Fig. 8 (right) for different altitudes at a given M0.78. As a result, the climb and take-off performance will be influenced by this sizing strategy. From this analysis, it could be concluded that the EM should be sized at relevant take-off conditions such that take-off and initial climb performance requirements are met. But in fact, the critical sizing case of the EM will depend on the level of HP,USE. For low values of HP,USE, due to the fact that the GTF still delivers a large proportion of the thrust at take-off, the sizing case of the EM will be at TOC conditions. The thrust available at take-off from the GTF is an outcome of the sizing of the GTF at TOC. It depends on the thrust lapse characteristics of the GTF and on the temperature limitation setting at take-off. Several de-rating scenarios can be envisaged to improve the life-cycle of the GTF according to the take-off requirements. The de-rating of the GTF will need to be adapted according to the EF maximal thrust available at take-off (which depends on the level of HP,USE, the resulting installed EM according to the TOC and the thrust lapse characteristics of the EF) to conserve the same take-off field length performance.

At higher levels of HP,USE, the EF plays a larger role in providing the required thrust during take-off and initial climb. Consequently the EM will have to be sized such that the DF can deliver the required thrust to fulfill the take-off and initial climb requirements. In this scheme, due to the impact of the EM power on the weight of the electrical system and its effects on the overall performance of the aircraft, the sizing of the EM would need to be thought as an optimization scheme whose objective is the minimization of the weight of the electrical system while fulfilling the take-off and climb performance constraints. The critical sizing case of the EM depends consequently on HP,USE, the thrust lapse characteristics of the TF and the thrust and power lapse characteristics of the EF. The analysis of the several sizing cases of the EM and the respective level of HP,USE is provided in Section 4.4 in the context of the quad-fan concept. 3.4. Discussion about the model and its limitations In the proposed concept, the EFs are operated at their maximal thrust available during the mission. Operational phases covering taxi-in/out, descent, landing and hold have been excluded, which are performed only with the TFs, and this was decided in order to maximize the utilization of the electrical system, and as a result increase largely the overall propulsion system efficiency. The maximal thrust provided by the EF is a function of the sizing and thrust lapse characteristics of the DF and of the maximum power of the EM installed. The difference in thrust requirement between TOC and cruise as well as the reduction in aircraft Gross Weight (GW) result in a necessity to throttle back the GTFs during cruise. The amount of thrust to be reduced is determined by the maximal thrust delivered by the EFs, the sizing of the GTF and the thrust required during cruise. Due to the throttling of the GTFs during cruise, the GTF can be subjected to operation within part-load conditions. The GTF can run into part-load conditions provided the HP,USE is sufficiently large (see Section 4.1). However, according to the sizing procedure of the GTFs and EFs at TOC (see Section 3.1), the amount of thrust required by the GTFs during cruise will mainly remain within the TSFC bucket of the GTF (see Fig. 7). In other words, the efficiency of the GTF will not be strongly impaired during cruise, which will benefit the overall system efficiency of the hybrid-electric motive power system. Only at large values of HP,USE the GTF operations into part-load might result into slight degradation of its efficiency. The efficiency change of the GTF with increasing HP,USE is discussed in Section 4.1.

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performance of the quad-fan hybrid-electric aircraft, an underwing podded, twin-engined reference aircraft with advanced gasturbine only was modeled for comparison purposes. Both aircraft reflect the same estimated technology freeze year of 2030 (YEIS 2035) and were sized for exactly the same set of Aircraft Top-Level Requirements. The advanced gas-turbine aircraft is referred as the “reference aircraft” in this technical article. The aircraft top-level requirements, the design mission definition and the aircraft sizing are described in the Sections 4.2 and 4.3 and IV.C respectively. 4.1. Parametric descriptors and qualitative pre-design investigation

Fig. 9. Electrical propulsion system architecture.

According to the selected operation of the EFs, the increase in HP,USE is limited to a value that would result in a shut-down of the GTFs during cruise due to fact that the EFs can deliver the required thrust alone. Indeed, in this situation, the EFs should not run at the maximal available thrust – the thrust should be adjusted to the cruise required thrust. In the investigation presented in this technical article, a maximal HP,USE value was set at 0.60 due to the nature of the implemented GTF model. Above this HP,USE threshold the design thrust of the GTF would have violated the validity of the implemented models and as a result of the low design thrust, it was judged that the selected GTF architecture would not be relevant anymore. 3.5. Electrical system characteristics A identical electrical propulsion system architecture as published by Pornet et al. [21] was selected and is represented in Fig. 9. It is composed of battery packs controlled by Battery Controller Units (BCUs) that deliver electric power to the High Temperature Super-conducting (HTS) EM equipped with a controller. Solid State Power Controllers (SSPC) protect the complete system from failure cases. The assumptions made in terms of components technology level are summarized in Table 1. The implemented allelectric sub-system reflects the same architecture as proposed in Pornet et al. [21]

4. Aircraft sizing and integrated performance The hybrid-electric propulsion system comprises two GTFs and two EFs installed in an underwing podded arrangement, and is integrated on a 180 PAX narrow-body transport aircraft. The sizing of the hybrid-electric propulsion system follows the scheme described in the previous section. In order to assess the integrated Table 1 Overview of propulsion components and technology performance assumptions. Component

Mass sizing parameter

Efficiency [%]

Source

Battery

1000–1500 Wh/kg

85.0–95.0

HTS-Motor Controller Converter HTS Cable SSPC

20.0 kW/kg 20.0 kW/kg 18.0 kW/kg 9.2 kg/m 44.0 kW/kg

99.5 99.5 99.5  100 99.5

Stückl et al. [46] Kuhn et al. [51] Brown [52] Brown [52] Brown [52] Xin [53] Brown [52]

An algebraic basis for the quantification of HP and the DoH for Energy (HE) was established by Isikveren et al. [19], and a correlation between HP and HE was suggested using the format of the Ragone diagram [20]. As was previously stated in the introductory remarks, Isikveren et al. [19] presented advanced aircraft tradestudy work using the parametric descriptor of HP,INS, which was found to be solely a function of the Supplied Power Ratio (Φ, related to converted power afforded by each energy carrier). In contrast, HE was found to be a more complex synthetic function described by comingling of Φ and the Activation Ratio (ϕ, describing the relative nature of utilization with respect to time afforded by the motive power device associated with each energy source). Furthermore, it was reasoned by Isikveren et al. [19] that aircraft design range exhibits a monotonic functional sensitivity with Activation Ratio, thus ϕ can be considered to serve as a proxy for design range. Apart from defining the two fundamental independent non-dimensional variables of Φ and ϕ for purposes of sizing and optimizing hybrid-electric aircraft propulsion systems, sensitivity studies involving variation of Φ and ϕ were also conducted by Isikveren et al. [19] An interesting observation pertains to the possibility of examining a series of hybrid-electric aircraft propulsion systems that deliver the same relative block fuel reduction or Energy Specific Air Range (ESAR, see Section 4.5) outcome. Generally, for a given gravimetric energy density and set of conversion, transmission and propulsive efficiencies, a constant relative block fuel reduction can be achieved with concurrent changes in Φ and ϕ. As a simple illustration of how the inter-play between Φ and ϕ could be manipulated for desired effect, the socalled HP,INS–HE Onion Curves chart given by Isikveren et al. [19] has been reconfigured as an HP,USE–HE correlation in Fig. 10. Each of the scenarios labeled ➊ through ➍ in Fig. 10 represent strategies and corresponding outcomes when in a generic sense one attempts to maintain the same relative block fuel reduction assuming a quad-fan morphology versus that of an original hybridelectric twin-fan configuration. The option as shown by ➊ considers keeping HE constant. Since ϕ (and design range, R) decreases, Φ (and subsequently η) needs to be increased in order to maintain the original value of HE. Although ϕ decreases, a magnified increase in Φ leads to a higher increased Gross Weight (labelled as IGW) as well as a corresponding improvement in lift-to-drag ratio (L/D) since constant wing loading during resizing means that the relative share of wing-to-fuselage zero-lift drag becomes favorable. Owing to a magnified increase in η coupled with a modest increase in IGW (and L/D), the vehicular efficiency metric, ESAR, improves in a magnified manner. Although this option is a much more energetically efficient strategy, the reduction in design range capability makes for a prohibitive choice. Option ➋ considers an increase in Φ (and subsequently η) whilst maintaining constant ϕ (and design range). Again, modest increase in IGW (and L/D) is indicative of this choice. The outcome for ESAR is a minor-tomodest improvement, and thus can be considered to be a viable strategy provided a design range increase is not predicated. A significant design range increase can be afforded using the option

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Fig. 10. Strategies and corresponding outcomes when maintaining the same relative block fuel reduction assuming a quad-fan morphology versus a twin-fan configuration.

➌ approach. In this instance, increases in together with a corresponding significant increase in IGW (and L/D) leads to a modest reduction in ESAR. Although there is likely to exists a moderate reduction in vehicular efficiency with this strategy, an increase in design range makes for a worthwhile possibility. Finally, as depicted with ➍ in order to keep HP,USE constant Φ (and subsequently η) need to be kept fixed and a magnified increase in ϕ (and design range) needs to occur. This leads to a magnified increase in IGW (and L/D) and the collective influence generates the least desirable ESAR result. Irrespective of the significant increase in design range this scenario would be considered not so desirable due to excessive degradation in vehicular efficiency. This simplified investigation and associated strategy review as shown in Fig. 10 was conducted in order to help formulate strategies for the more sophisticated trade-studies presented in this technical article. 4.2. Aircraft Top-level Requirements The following Aircraft Top-Level Requirements (ATLeRs) were stipulated for the design of the conventional and hybrid-electric aircraft. The YEIS was set as 2035. The design payload is 180 PAX

assuming 102 kg per PAX. Compliance with the airworthiness regulations CS-25 and FAR 25 transport category were administered, and it is worth noting that the minimum second segment climb gradient requirement in this study conformed to one applicable to four-engined aircraft. Field performance is characterized by a Take-Off Field Length (TOFL) at ISA, sea-level not being greater than 2200 m, and an approach speed less than 145 KCAS. Additionally, the time-to-climb to initial cruise altitude is limited to 25 min. 4.3. Design mission The defined mission profile consists of a taxi-out, take-off at sea level and ISAþ10 °C, climb at ISAþ10 °C with speed schedule 250 KCAS/300 KCAS/M0.76 until an initial cruise altitude at FL350. The cruise, performed at M0.76 and ISAþ10 °C, is followed by a mirrored descent, landing and taxi-in. Reserves and contingency fuel is in accordance with EU-OPS 1.255, namely, 5% trip fuel contingency cruise, 30 min hold at 1500 ft and 100 nm alternate. However, due to the combined use of the GTFs and EFs a contingency cruise account based on trip fuel is not relevant anymore.

Fig. 11. Power profiles of the hybrid-electric propulsion system covering block and reserves/contingency operational phases.

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 Fig. 12. Relative change in block fuel versus relative change in block ESAR.

EM. To gain an insight into this sizing criterion, the performance of the reference aircraft was analyzed. By sizing the GTFs at TOC and according to the implemented GTF model, the TOFL (ISA, SL) without de-rating is below 1900 m with time-to-climb at around 20 min. (for each of the investigated design ranges). The take-off and climb performance, outcome of the sizing of the GTF, exceed consequently the expectations of the ATLeRs. As a result, within the aforementioned HP,USE range and recalling the effects of sizing the EM at TOC (refer to Section 3.3), the TOFL and time-to-climb can be let increasing compared to the nonrated reference aircraft while still remaining within the ATLeRs. When HP,USE reaches the threshold of 0.25, the time-to-climb constraint limitation of 25 min is reached. For HP,USE values between 0.25 and 0.50, time-to-climb was found to be the critical sizing case for the EM installed power. Consequently the EM power is sized such that time-to-climb does not exceed 25 min. For HP,USE values above 0.50, TOFL (ISA, SL) turns out to be the sizing condition of the EM. The EM maximal power is consequently sized such that the TOFL remains within the limit of 2200 m.

The contingency cruise reserve condition was adapted and expressed as an equivalent time criterion defined as 10% of the block time. As mentioned earlier, taxi-in/out, descent, landing and hold operations are performed with the GTFs only (no additional windmilling drag component was assumed in this analysis due to shutdown of the EFs during these segments). Fig. 11 illustrates the power profiles of both the GTFs and EFs over the course of the complete block and reserves/contingency operational phases.

These sizing cases are the underlying EM sizing criterion in each of the results presented in the following sections. An interesting note is that the critical sizing scenarios remain almost identical for each of the design ranges investigated. This fact is to be explained and discussed in the following sections of this technical article.

4.4. Aircraft sizing

The relative change in block Energy Specific Air Range (ESAR with units nm/kWh) [55], which is a vehicular efficiency figure-ofmerit, against the relative change in block fuel assuming the reference aircraft as a datum are illustrated in Fig. 12 according to variation of HP,USE and for an interval of design ranges between 900 nm and 2100 nm. The first observation from results indicate that within the investigated design space the vehicular efficiency of all hybridelectric concepts are impaired compared to the reference aircraft. The relative change in block ESAR which characterizes the distance traveled per total energy consumed, for the given set of assumptions here, is indeed always negative. The analysis of the relative change in block fuel shows that for low HP,USE the block fuel consumption is larger compared to the reference aircraft for every investigated design range. By increasing HP,USE block fuel reduction can be achieved whose amount depends upon the design range, but at the penalty of a further reduction in vehicular efficiency. These trends can be explained by the following design considerations:

4.4.1. Sizing guidelines For the sizing of the aircraft, the following design rules and assumptions were employed. The wing is sized according to a constant wing loading of 645 kg/m² (to avoid 1.3 g buffet onset limitations) and a constant wing aspect ratio of 12.5 (reflecting the YEIS 2035 technology standard). As the fuselage geometry is kept fixed, the underfloor volume represents a constraint for the accommodation of batteries and for the storage conventional cargo. The baggage volume for the advanced gas-turbine only aircraft is of 0.22 m³ per PAX. The volume constraint problem has been analyzed assuming a linear variation of the battery volumetric specific energy with respect to the gravimetric specific energy according to a factor of 1000 kg/m³ which includes the volume of the battery and of the thermal management system. For instance, for a battery gravimetric specific energy of 1.5 kWh/kg, the battery volumetric specific energy is assumed to be 1500 kWh/m³. It is also assumed that aircraft loadability due to battery installation would be achieved in a more detailed study. Generally, this value of 1000 kg/m³ gravimetric-to-volumetric specific energy is rather conservative when compared to 2000–3000 kg/m³ indicative of contemporary terrestrial hybrid vehicles [54]. Nonetheless, for an initial assessment involving the application of high performance batteries and associated thermal management in the context of aeronautical application the authors feel it is an appropriate target. 4.4.2. Electric motor critical sizing conditions The sizing of the hybrid-electric propulsion system and in particular of the EFs was generally discussed in Section 3.3. The critical sizing scenarios of the EM are elaborated in the context of the quadfan concept investigated as function of HP,USE in the following:

 For HP,USE values below 0.25, the DF power required at TOC condition was found to be the critical sizing condition of the

4.5. Relative change in block esar versus relative change in block fuel

1. Compared to the twin-engined reference aircraft, the installation of EFs and the associated electrical system adds weight and results in an aerodynamic penalty due to the additional wettedarea of the nacelles. At low values of HP,USE the slight increase in overall propulsion system efficiency (provided by the electrical system and its highly efficient components) does not counteract these negative aspects entirely. Consequently, the quad-fan aircraft displays higher fuel burn consumption and a lower vehicular efficiency compared to the reference aircraft. 2. By increasing HP,USE the utilization of electrical power increases, which enables a larger improvement to the overall propulsion system efficiency. Due to the use of electrical energy and by virtue of an increase in overall propulsion system efficiency, the block fuel can be significantly reduced. However, because of the additional weight of the electrical system and of the batteries,

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reference aircraft. As a result, cargo volume represents a volumetric constraint for the housing of batteries and cargo. A typical cargo volume of 0.14 m³ per PAX is displayed (as solid triangle symbols) in Fig. 12, taken to be appropriate for regional market segments, provides an indication of the extent restrictions cargo volume limitations impose. At a design range of 900 nm, the cargo volume limit of 0.14 m³ per PAX is reached at an HP,USE of 0.47. By increasing the design range, the amount of batteries required becomes larger. As a result, the cargo volume limit is reached at lower HP,USE values. For instance, considering a design range of 1700 nm, the cargo volume per PAX limitation is reached at an HP,USE of 0.25. In summary, implications of the hybrid-electric propulsion system at aircraft level with respect to an increase in HP,USE is characterized by:

 An increase in the amount of electric energy used, which results in an increase in battery mass;

 An increase in aircraft weight (due to increasing required batFig. 13. Relative change in block fuel versus relative change in block COSAR according to a specific fuel price of 6.00 USD/USG and a specific electricity price of 0.07 USD/kWh.

 

tery mass and the sizing of the electrical system) which initiates ever more significant sizing cascade effects; An enhancement to the overall propulsion system efficiency due to more extensive use of electrical energy, with a corresponding reduction in block fuel consumption; and, An overall decrease in vehicular efficiency due to weights effect, which tends to overwhelm the benefit of the hybrid-electric system and L/D improvement due to all-axes wing scaling.

4.6. Relative change in block COSAR versus relative change in block fuel

Fig. 14. Relative change in MTOW and relative change in overall propulsion system efficiency (HP,USE as a proxy) at for an interval of design ranges.

the MTOW of the aircraft is noticeably increased (see also Fig. 14). Even figuring in an improvement to L/D as an outcome to an all-axes scaling of the wing (constant wing loading and fixed fuselage dimensions), ESAR is still degraded compared to the reference aircraft. For design ranges below 1300 nm, block fuel reductions beyond 35% are achieved while the block ESAR is reduced to values between  5% and 20%. Between 1300 nm and 1700 nm, the reduction in block fuel achievable is between  20% and  40%. The degradation in ESAR becomes, however, significant with values between  20% and  30%. 3. Increasing the design range results in a larger energy requirement, and notably, an increase in the amount of batteries required, which triggers significant sizing cascade effects. Although it still leads to a decrease in potential fuel burn reduction the degradation in ESAR becomes much larger. For design ranges above 1900 nm, no significant improvement in fuel burn compared to the reference aircraft is obtained with increasing HP,USE. For this preliminary investigation of quad-fan aircraft concepts, the entire fuselage geometry was kept identical to the one of the

In order to bridge the gap between a purely technical investigation and conducting an economics analysis of hybrid-energy aircraft concepts, the cost-based figure-of-merit COSAR (Cost Specific Air Range) was introduced by Pornet et al. [56]. The relative change in COSAR is presented in Fig. 13 as an integrated block value. COSAR with units [nm/USD] characterizes the distance travel per cost of energy. COSAR depends upon the respective specific cost of the energies considered. In this assessment, the specific fuel cost is set at 6.00 USD/USG and the specific cost of electricity is set at 0.07 USD/kWh [56]. In other words, the specific cost of electricity represents in this evaluation around 40% of the specific cost of the fuel. Due to the lower specific price of electrical energy compared to fuel, reducing fuel consumption versus utilizing electrical energy becomes more advantageous. This is explained by the fact that the change in block COSAR compared to the reference aircraft becomes positive at high HP,USE for design ranges between 900 nm and 1500 nm. For an HP,USE of 0.60, the block COSAR can be increased from 10% to 30% for design ranges 1300 nm and 900 nm respectively. For longer design ranges, the large increase in required energy indicated by the significant degradation in ESAR (see Fig. 13) and in particular the increase in required electric energy overcomes the benefit obtained from the lower specific price of electricity. As a result, for design ranges above 1500 nm, no improvement in block COSAR is achieved. At a design range of 1700 nm and an HP,USE of 0.55, the block COSAR is 7% less than the reference aircraft. Comparing statements drawn from the analysis of ESAR variation (Fig. 12) and of COSAR change (Fig. 13) underlines the importance of considering economics analysis in order to determine a “best and balance” hybrid-electric concept. 4.7. Relative change in MTOW versus the Degree-of-Hybridization for Useful Power Weight effects during sizing of the installation electrical system and of the batteries are reflected by the relative change in MTOW

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indicated in Fig. 14, which displays significant increase in MTOW against the reference aircraft. At a design range of 900 nm and an HP,USE of 0.60, the MTOW of the hybrid-electric aircraft is 49% larger than the reference aircraft and 85% larger for 1300 nm. This weight effect counteracts strongly the benefit of increased overall propulsion system efficiency provided by the highly efficient electrical components as demonstrated previously with the evaluation of relative change in ESAR (see Fig. 12). It is interesting to highlight the change of paradigm in aircraft design when designing and sizing a fuel-battery hybrid-electric aircraft. First of all, fuel burn reduction can be achieved while the MTOW of the aircraft is increased. For a single energy source aircraft, assuming a given aerodynamic and propulsion system efficiency, a decrease in block fuel can only be achieved through a reduction in aircraft mass. In the case of fuel-battery hybrid-electric aircraft, block fuel reduction can be achieved while the mass of the aircraft increases. This apparent paradox results from the utilization of another source of energy, namely electrical energy, and from the increase in overall propulsion efficiency. Secondly, a reduction in fuel burn reduction does not mean automatically an improvement in vehicular efficiency. When the change in weight is larger than the change in overall propulsion system efficiency, the total energy consumed (fuel and electrical energy) increases (see relative change in ESAR in Fig. 12). The aircraft utilizes energy in a less efficient manner, therefore its vehicular efficiency decreases while a significant fuel burn reduction could still be achieved. 4.8. Relative change in point ESAR versus the Degree-of-Hybridization for Useful Power To gain more insight into the contributions affecting the change in vehicular efficiency, the relative change in ESAR compared to a reference is expressed as

ΔESAR =

Δη⋅Δ(L/D) ΔW

(3)

As can be gauged in Eq. (3) constituent parameters of overall propulsion system efficiency (η), L/D and Gross Weight (W) all combine to define the ESAR qualities of a given hybrid-electric aircraft concept. Variation of these constituent parameters is shown in Fig. 15 corresponding to a cruise start point (ISAþ10 °C, FL350, M0.76) for an aircraft sized for design range of 1300 nm. Focusing first on the L/D, the installation of the EFs results in

additional wetted area due the extra nacelles and pylons, however, this tends to reduce the size of the GTFs. As a result, at an HP,USE of 0.05, the L/D is roughly the same as that of the reference aircraft. Due to sizing effects, the L/D increases with HP,USE; for instance, at an HP,USE of 0.60, the change in L/D is increased by 9%. With the introduction of the more efficient electrical system, the overall propulsion system efficiency is increased, and this is indicated by a growing HP,USE. At an HP,USE of 0.05, Δη is around 3% and is increased up to 46% at an HP,USE of 0.60. The positive influence of the increase in L/D and η is countervailed by the significant increase in W. The installation of the EFs, its associated electrical system, battery weight and the resulting cascading effects lead to 5% higher gross-weight at an HP,USE of 0.05 and 88% larger at an HP,USE of 0.60. These combined changes in L/D, η and W lead to the following change in ESAR at cruise start: it remains negative in the investigated range of HP,USE and remains approximately constant at  2% up to an HP,USE of 25%. Above this value, the contribution of MTOW change overwhelms the change in L/D and η leading to a net reduction in the ΔESAR down to 14% at an HP,USE of 0.60. The change is represented here for a single point within the mission; however, it provides a good insight into the design aspects which contribute to changes in ESAR. By integrating these changes along the segments that constitute the block and diversion legs, the integrated value of block ESAR change presented in Fig. 12 is obtained. 4.9. Influence of Degree-of-Hybridization for Useful Power on the combustion-based propulsion system efficiency To analyze the influence of the EFs running at maximal available thrust during cruise on the efficiency of the GTFs, the thrust lapse of the GTF between the thrust required at TOC and the thrust required at start cruise condition (ISAþ10 °C, FL350, M0.76) are illustrated in Fig. 16. The efficiency of the conventional propulsion system is represented at start, mid-cruise and end-cruise conditions with respect to HP,USE. Recalling the TSFC bucket of the GTFs spans 70%-100% of the maximum thrust (see Fig. 7), it can be concluded that for HP,USE values up to 0.45, the GTF still operates well within the TSFC bucket. As a result the efficiency of the GTF even at mid-cruise and end-cruise remains almost unchanged. However, for HP,USE values above 0.45, the thrust lapse increases as the thrust delivered by the EFs during cruise increases. Consequently, the efficiency of the GTF is slightly impaired. In Fig. 7, the 100

100 Change in L/D Change in gross−weight Change in propulsion system efficiency Change in ESAR

90 80

90 85

Study Settings: Point performance: start cruise ISA+10°C, FL 350, Mach 0.76 Design range1300 n.mi ebattery = 1.5kWh/kg

60 50 40

Value in percent [%]

Relative change [%]

70

30 20 10

80 75 70 Study Settings: Point performance: top−of−climb ISA+10°C, FL 350, Mach 0.76 Point performance: cruise ISA+10°C, FL 350, Mach 0.76 Design range1300 n.mi ebattery = 1.5kWh/kg

65 60 55 50

0

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GTF thrust lapse Cruise/Top−of−climb Conventional system efficiency start cruise Conventional system efficiency mid cruise Conventional system efficiency end cruise

95

0

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USE

50 [%]

60

Fig. 15. Point performance analysis at cruise start (ISAþ 10 °C, FL350, M0.76) for a design range of 1300 nm.

35

0

10 20 30 40 Degree of hybridization for power, Hp

USE

50 [%]

60

Fig. 16. Analysis of the combustion-based propulsion system efficiency variation during cruise (ISAþ10 °C, FL350, M0.76) for a design range of 1300 nm.

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4.10. Influence of the Degree-of-Hybridization for Useful Power on take-off performance Take-off performance was analyzed according to assumption the One Engine Inoperative (OEI) case results only from a failure of the critical GTF only, whereas the EFs are still operating according to the useful EM power and the off-design characteristics of the DF. As a matter of interest, it can be noted that according to the assumption for the theoretical case of HP,USE equals 1.0, the All-Engines Operational (AEO) and OEI performance would be identical. A detailed study about the influence of different failure modes of the hybrid-electric propulsion system during take-off operations is subject to future work. Notwithstanding this, the current assumption could be legitimized by the knowledge that there would exist a greater level of reliability using electrical components. Because of the assumption that the EFs are unlikely to fail compared to the GTFs, take-off performance is improved with increasing HP,USE. Indeed, as the TOFL and second segment climb gradient are generally driven by an OEI case, increasing the thrust contribution of the EFs during take-off compared to the GTFs enhances the OEI takeoff performance. The effect at aircraft level is visualized by plotting the evolution of the AEO and OEI thrust-to-weight ratio, (T/W)AEO and (T/W)OEI, respectively at a given take-off condition (ISA, SL, M0.20) in Fig. 17. The variation of the ratio between TOFL and the TOFL limitation of 2200 m stipulated by the ATLeRs reflects the scenario employed for the sizing of the EMs as discussed in Section 4.4. As TOFL is not a critical sizing criterion up to an HP,USE of 0.50, it is an outcome of the overall sizing process. According to other active sizing criteria TOFL increases up until the TOFL requirement defined in the ATLeRs is 110 100 90

Value in percent [%]

80 70 60 50 40

Study Settings:

TOFL/TOFL

Point performance: take−off ISA, SL, Mach 0.2 Design range1300 n.mi e = 1.5kWh/kg

T/W

0 1300

10 15

−5

20 30 Design Range [nm]

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40

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45 Study settings: ebattery = 1.0kWh/kg −40

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−25

−20

50

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0

Relative change in block ESAR [%] Fig. 18. Relative change in block fuel versus relative change in block ESAR.

reached. Consequently, (T/W)AEO decreases with increasing HP,USE up to 0.50. At an HP,USE value of around 0.50, TOFL becomes the critical sizing criterion. However, it can be observed that instead of remaining constant, (T/W)AEO still slightly decreases. Besides the change in aerodynamic efficiency due to sizing effect with increasing HP,USE, it is also a result of increases in (T/W)OEI with higher HP,USE. In other words, by increasing HP,USE more thrust per weight is available during the OEI case. 4.11. Sensitivity study according to battery gravimetric specific energy The assumptions made in terms of battery gravimetric specific energy have significant implications on the presented results. Consequently, the sensitivity with regards to the battery gravimetric specific energy was analyzed. The relative change in block fuel and block ESAR is represented in Fig. 18 assuming a battery gravimetric specific energy of 1.0 kWh/kg at cell-level. It can be first observed that only for design ranges strictly below 1500 nm block fuel reductions can be achieved against the projected, combustion-based only aircraft. For a given range and a given HP,USE, due to the higher resulting weight of the batteries and corresponding sizing cascade effect the potential in block fuel is reduced compared to results assuming 1.5 kWh/kg. Moreover, for the same design range and HP,USE the weight impact results in a larger degradation in block ESAR compared to one calculated at 1.5 kWh/kg. In addition, for a given design range according to the assumption made in terms of the gravimetric specific energy, the volumetric constraint for the housing of the battery occurs at a lower HP,USE.

5. Influence of the Degree-of-Hybridization for Useful Power on flight technique optimization

T/WOEI

20 10 10

1500

5

ATLeRs

battery

0

Cargo volume/PAX = 0.14m³

AEO

30

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Relative change in block fuel [%]

TSFC of the GTF is reduced by roughly 4% at a thrust setting of 60% of maximum thrust. When computing the change in efficiency of the combustion-based system between, for instance, HP,USE of 0.30 and 0.60, a similar value of about  4% is obtained. Moreover, due to the fuel consumed during cruise and the resulting decrease in aircraft mass, the required thrust reduces from the start to the end of the cruise phase. As the result, the thrust lapse incrementally increases at mid-cruise and end-cruise leading to a more pronounced part-load operating GTF. However, this effect is counterbalanced by a corresponding reduction in fuel flow with increasing HP,USE, thus the change in aircraft mass becomes less pronounced during cruise.

17

20

30

40

50

Degree of hybridization for power, Hp

USE

60

[%]

Fig. 17. Takeoff Field Length, thrust-to-weight ratios of All-Engine Operational and One Engine Inoperative conditions against Degree-of-Hybridization for Useful Power, design range of 1300 nm.

An assessment of the influence HP,USE has on flight technique optimality together with implications to operational procedures is reviewed in this section. For a design range of 1300 nm, the ESAR altitude-speed sensitivity is investigated at a GW of 98% MTOW for increasing HP,USE. The results of this sensitivity study are illustrated in Fig. 19 (left) for an HP,USE of 0.25 and in Fig. 19 (right) for an HP,USE of 0.40. A contrast of these results can be conducted when surveying results of a projected, combustion-based only reference aircraft displayed in Fig. 21 within Section 4.2. In addition, Fig. 22 provides flight technique sensitivity

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Fig. 19. ESAR altitude-speed sensitivity versus the Degree-of-Hybridization for Useful Power for a design range of 1300 nm and at 98% of MTOW; (left) Degree-of-Hybridization for Useful Power of 0.25; (right) Degree-of-Hybridization for Useful Power of 0.40.

Fig. 20. Geometrical and dimensional changes in accordance with the Degree-ofHybridization for Useful Power.

information for a tentatively chosen benchmark hybrid-electric aircraft candidate with attributes of design range 1300 nm and HP,USE equal to 0.30 (as detailed in Section 6). In all the above-mentioned charts associated with hybridelectric aircraft, the change in ESAR optimum indicates that with increasing HP,USE, the optimum flight technique tends towards lower speed and lower altitude. At an HP,USE of 0.25, the maximum ESAR was found to be M0.70/FL345 [Fig. 19 (left)]. At an HP,USE of 0.30 the optimum is at M0.69/FL335 (Fig. 19), whereas at an HP,USE of 0.40 the ESAR optimum occurs at M0.67/FL310 [Fig. 19 (right)]. A similar trend was already identified during the investigation of an universally-electric aircraft concept [47] and the comparison of its flight technique optimality against a projected gas-turbine only aircraft [55]. The explanation of this change in altitude-speed optimality is rather complex as it involves the sensitivity of the overall propulsion system efficiency and the sensitivity of the aerodynamic efficiency with speed and altitude. For an increasing HP,USE, flight technique optimality of the quad-fan aircraft is driven more importantly by the characteristics of the EFs as they produce a larger proportion of the thrust. The altitude-speed condition has less influence on the EF characteristics compared to that of the GTF characteristics. Indeed, in this analysis the EMs are assumed to be thermally managed, and consequently, the efficiency is assumed to

Fig. 21. Reference aircraft; (left) Altitude-speed sensitivity for a design range of 1300 nm and a gross-weight of 98% of MTOW; (right) Speed sensitivity versus gross-weight variation at a design range of 1300 nm, FL350 and a specific fuel price of 6.00 USD/USG.

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Fig. 22. Hybrid-electric aircraft – (left) Altitude-speed sensitivity for a design range of 1300 nm, Degree-of-Hybridization for Useful Power of 0.30 and gross weight 98% of MTOW – (right) Speed sensitivity versus gross weight variation for a design range of 1300 nm, Degree-of-Hybridization for Useful Power of 0.30, FL350, specific fuel price of 6.00 USD/USG and variations in specific electricity price.

be constant with altitude-speed. A similar assumption is made for the other components that constitute the entire advanced electrical architecture. In addition, as this analysis was conducted on a point performance basis, it is acceptable to assume constant battery efficiency. Consequently, the change in the electrical system efficiency is only driven by the change in propulsive efficiency of the DF with altitude-speed variation. In contrast, the efficiency of the GTF is strongly influenced by the effect of altitude-speed variation on the gas-turbine characteristics. The change in overall propulsion efficiency with altitude-speed becomes consequently less pronounced HP,USE increases. As a result, by increasing HP,USE ESAR optimality is mainly driven by the sensitivity of the aerodynamic polar characteristics of the aircraft goverened by speed and altitude. These characteristics tend notably to slow down the optimum speed of the aircraft, and commensurate with this the optimum tends towards a lower flight altitude as well. Following the philosophy of the Long Range Cruise (LRC) definition (characterized by 99% SAR maximum), altitude-speed techniques of M0.72/FL350 for HP,USE of 0.25, and, M0.71/FL350 for HP,USE of 0.40 could be defined at 99% of ESAR maximum. Such a flight technique is denoted in this paper by the term Long ESAR Range Cruise (LERC). In view of these results, it can be concluded that the effect of HP,USE on flight technique optimization will not drastically change the altitudespeed operation of hybrid-electric aircraft, and that they could be integrated without significant changes into the existing commercial transport network routes of current turbofan airliners.

6. Aircraft characteristics and benchmark A design range of 1300 nm was selected for the benchmark of the hybrid-electric. The selection is justified by in-house projected market analysis for YEIS 2035, which indicates that this design range covers 90% of the projected cumulative stage lengths for the narrow-body class of aircraft [57, 58]. Recalling Fig. 12, since the housing of batteries is volumetrically constrained at an HP,USE above 0.33 a design point with an HP,USE of 0.30 was consequently selected for the benchmark analysis to follow.

Fig. 14. According to the sizing setting of the wing with constant aspect ratio and constant wing-loading, the wing area is increased by 28% and the wing span by 13%. As a consequence, with a 39.1 m wing span the ICAO Code C airport compatibility is violated and needs to be reclassified as Code D. As a result of the sizing effect, aerodynamic efficiency, L/D, is increased by 5%. According to the downsizing of GTFs with the introduction of EFs, design thrust has been reduced by 14%. It results in a modest increase in TSFC, þ 2% at design point. At the selected HP,USE of 0.30, the sizing criterion of the EMs is the time-to-climb (see Section 4.4) requirement. The resulting installed power of a single EM is 2220 kW. As illustrated in Fig. 17, TOFL is increased by 13% compared to the non-rated TOFL performance of the reference aircraft. To provide some indication to the reader of the geometrical and dimensional changes, an overlay of the hybrid-electric aircraft sized for HP,USE of 0.10, 0.30 and 0.50 is provided in Fig. 20. As readily seen, aircraft morphologies beyond that of DWB, or, a “tube-and-wing” combination does not appear to be an outcome. (Tables 2 and 3) 6.1.2. Aircraft mass breakdown As shown in Table 1 hybrid-electric aircraft sizing effects have produced an outcome where the structural weight increased by 17%. Focusing on the propulsion system, the 14% down-size in design thrust of the GTF leads to a decrease of 14% in the combustion-based power plant weight. However, due to the installation of the EFs, which includes the weight of the DFs and of the EMs, the combined propulsion system weight is increased by 24%. In addition, the installation of the electrical propulsion system Power Management and Distribution (PMAD) with 1348 kg results in a 37% increase in electrical system weight. Consequently, the Operational Empty Weight (OEW) of the hybrid-electric aircraft is 16% higher. The release fuel (including block and reserves/contingency) is reduced by 8%. Recalling Fig. 12, the block fuel reduction was found to be around 15%. The battery mass installed is 11,740 kg. 6.2. Optimum flight technique analysis

6.1. Aircraft benckmark 6.1.1. Aircraft main data Comparison of the hybrid-electric concept with the reference aircraft (Table 1) indicates a 28% increase in MTOW as illustrated in

To gain more insights into the implications of the hybridelectric propulsion system on flight technique optimization, the optimum flight techniques of the benchmark hybrid-electric aircraft are compared against the reference aircraft in this section.

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Table 2 Aircraft main data comparison at a design range of 1300 nm and HP,USE of 0.30. Aircraft Main Data

Unit

Reference aircraft

Quad-fan hybrid

Δ Reference

MTOW MLW OEW/MTOW Design Payload/MTOW Release Fuel/MTOW Total Battery/MTOW Thrust-to-Weight ratio (SL, Static) Wing Loading Wing Ref. Area (Airbus Gross) Wing Aspect Ratio (Airbus Gross) Wing Span Wing Leading Edge Sweep GTF Design Thrust (ISA, FL350, M0.78) EF Design Thrust (ISA, FL350, M0.78) TSFC (ISA, FL350, M0.78) Electric Motor Installed Power L/D (initial cruise, ISA, FL350, M0.78, MTOW brakes release) TOFL (ISA, SL) Approach speed (ISA, SL, MLW)

[kg] [kg] Dimension Dimension Dimension Dimension Dimension [kg/m²] [m2] Dimension [m] [deg] [kN] [kN] [g/kNs] [kW] Dimension [m] [KTS]

60,840 59,010 0.594 0.302 0.104 NA 0.320 645 94.3 12.5 34.6 26.4 18.1 NA 13.29 NA 18.1 1850 142

77,730 76,170 0.538 0.236 0.075 0.150 0.290 645 120.4 12.5 39.1 26.4 15.6 6.8 13.50 2220 19.0 2080 143

28% 29%  9%  22%  28% NA  12% 0% 28% 0% 13% 0%  14% NA 2% NA 5% 13% þ0%

Table 3 Aircraft mass breakdown comparison at a design range of 1300 nm and HP,USE of 0.30. Mass Breakdown Summary Unit Reference aircraft

Quad-fan hybrid

Δ Reference

less

less

performance offered by Torenbeek [59] who had based his derivations on earlier work presented by Hale [60]. Torenbeek declares the Maximum Range Cruise (MRC), MMRC, speed as

MMRC =

Structure Propulsion System Conventional Powerplant Electric Motors ( þcontrollers) Ducted Fans

[kg] [kg] [kg]

18,440 4132 3861

21,487 5117 3553

17% 24%  14%

[kg]

NA

444

NA

NA

1119

NA

Equipment Furnishing Hydraulic System (antiicing) Electrical System PMAD propulsion Sub-systems Instruments Operational Items

[kg] [kg] [kg]

10,012 4435 173

11,640 4435 173

16% 0% 0%

[kg] [kg] [kg] [kg] [kg]

4460 NA 4460 944 3580

6088 1348 4740 944 3580

37% NA 6% 0% 0%

[kg] [kg] [kg]

36,159 18,360 6317

41,818 18,360 5806

16% 0%  8%

[kg] [kg]

NA 60,840

11,740 77,730

NA 28%

OEW Payload Release Fuel (blockþreserves) Total Battery MTOW

less less less less less

6.2.1. Reference aircraft The altitude-speed sensitivity of the reference aircraft is presented in Fig. 21. Due to the use of a single energy source, Specific Air Range (SAR), ESAR or COSAR figures-of-merit lead to the same sensitivity for the identification of flight technique optimality. At a GW condition of 98% MTOW, the optimum corresponds to M0.70/ FL350 in Fig. 21 (left). The LERC flight technique (defined as 99% ESAR maximum) would enable an increase in speed to M0.73 at FL350. Compared to a state-of-the-art narrow-body short-to-medium range aircraft, LRC was found to be slightly lower for a similar flight condition. The origin of this reduction lies in the extent of improvements in the overall propulsion system and aerodynamic efficiency provided by the advanced technology implemented in the reference aircraft and how they trade-off between one another when identifying an speed optimal condition. This outcome can be best explained via the theoretical treatment of optimum cruise

⎛ 2g ⎞ ⎛ W ⎞ ⎜ ⎟⎜ ⎟ ⎝ σ ⎠ ⎝ SW ⎠

4

⎛ dC /dC2 ⎞ ⎛ 2 + η ⎞ M ⎜ D L ⎟⎜ ⎟ ⎝ CDo ⎠ ⎝ 2 − ηM ⎠

(4)

where g is acceleration due to gravity, s is the density lapse ratio, SW is the reference wing area, dCD/dCL2 is the vortex-induced drag factor (including compressibility effects), CDo is the zero-lift drag coefficient (including compressibility effects), and ηM is the rate change in overall propulsion system efficiency with respect to Mach number. Now, assuming for a given altitude the Thrust Specific Power Consumption (TSPC, with units W/N) [55] varies with Mach number, for any type of power plant ηM can be obtained using logarithmic differentiation of the identity η¼Ma/TSPC, namely,

ηM =

⎛ M ⎞ dη d ln η =⎜ ⎟ d ln M ⎝ η ⎠ dM

⎛ M⎞ d ⎛ ⎛ M ⎞ d TSPC ⎤ TSPC ⎞ a ⎡ ⎟ ⎟ e (ln M − ln TSPC) = ⎜ M =⎜ ⎟a ⎢1 − ⎜ ⎥ ⎝ TSPC ⎠ dM ⎦ ⎝ Ma ⎠ TSPC ⎣ ⎝ η ⎠ dM =1−

M ⎛ d TSPC ⎞ ⎟ ⎜ TSPC ⎝ dM ⎠

(5)

with the parameter, a, denoting the local sonic velocity. Upon substitution of Eq. (5) into Eq. (4), owing to the transcendental (simple iteration) nature of the collective expression, small incremental changes in M and in dTSPC/dM are expected to occur even for [significant] variations in values of TSPC as nominated by the designer. The implication is in a simplified sense any reduction in TSPC would impart the greatest influence, thus leading to a reduction in ηM, thereafter generating a reduction in MRC speed. As can be gauged from inspection of Eq. (4) increasing wing loading (W/SW) and/or decreasing CDo will tend to increase the MRC speed. In contrast, if again substitution of Eq. (5) into Eq. (4) is considered, a decrease in dCD/dCL2 will tend to lower the MRC speed. Although the above construct given by Eq. (4) relates to quantifying changes in MRC speed, it is assumed the same amount of relative change due to variation in TSPC will approximately occur if an LRC speed convention is adopted, i.e. a fractional change in MLRC is approximately equal to the fractional change in MMRC. For a given flight altitude and ambient conditions, and varying aircraft instantaneous gross weight, the inference here is the rate change in SAR degradation with respect to Mach number, dSAR/dM, is approximately the same for LRC speed variation.

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6.2.2. Hybrid-electric aircraft The ESAR optimum of the hybrid-electric aircraft presented in Fig. 22 (left) shows a slight reduction in optimum speed and altitude with M0.69/FL335 compared to the reference aircraft. However, considering LERC, a flight technique of M0.72/FL350 can be defined which is closed to the one defined for the reference aircraft. In Fig. 22 (right), speed sensitivity versus GW variation is represented for MRC, the Maximum ESAR Range Cruise (MERC), the Maximum COSAR Range Cruise (MCRC) and LRC. The ESAR metric takes into account the total energy (the combined energy of fuel and electrical energy), while in the SAR metric only the fuel energy is considered. As indicated in Section 3.4, the EFs are operated at their maximum thrust during cruise while the GTF throttle can vary. As a result, the electrical energy is mainly a function of the flight time, or in other words, the flight speed. By flying faster, the cruise time is decreased for a given stage length, reducing the electrical energy required during cruise. Consequently, the minimum energy condition can be reached at higher speeds than for the condition for maximum SAR. The optimum speed is, however, counter-balanced by the higher fuel consumption at higher speeds. The change in optimum speed between SAR and ESAR was found to be small. Due to the investigation of the COSAR metric, the results depend upon the relative specific price of each energy source. In this application, the fuel price was set at a constant 6.00 USD/USG, whereas the electricity price was varied assuming 0.07 USD/kWh and 0.16 USD/ kWh. With a specific electricity price of 0.07 USD/kWh, which represents about 40% of the investigated specific fuel price, the fuel energy price has a higher contribution to the COSAR metric. Consequently, the COSAR figure-of-merit tends to an optimum speed slower than that of the ESAR optimum, but still slightly faster than that of the SAR optimum. When the specific electricity price equals the specific fuel price, which is the case at 0.16 USD/kWh, the COSAR speed optimum coincides with the ESAR optimum as both energy contributions are equally weighted in the COSAR metric. Compared against the reference aircraft whose results are presented in Fig. 21 (right), it can be noticed that the change in speed sensitivity is not altogether different using either the MERC or MCRC speeds. In terms of absolute COSAR value expressed in [nm/USD], a low electricity price leads to large improvement in the COSAR metric as it was already highlighted in Section 4.6. However, it is important to highlight that in this particular study, COSAR results are representative of an instantaneous cruise point calculation, whereas in Fig. 13 the COSAR value is integrated along the entire block mission (which includes all the transient flight phases). This explains why the changes in block COSAR are less than the change in COSAR, which can be calculated according to Fig. 21 (right) and Fig. 22 (right).

7. Conclusion Consideration given to the application of hybrid-electric propulsion systems for transport aircraft has dramatically expanded the design space in aircraft design. Because of the decoupling between the energy generation and the Useful Power production, the selection of the optimum energy system, the optimum integration of the propulsors as well as the optimum operation of the overall hybrid-electrical propulsion system need to be determined according to the aircraft requirements. In this contribution, the implication of an increasing Degree-of-Hybridization for Useful Power (HP,USE) on the design of a fuelbattery narrow-body transport aircraft was assessed. In order to freely vary the HP,USE, a quad-fan aircraft concept equipped with two advanced Geared-Turbofans (GTFs) and two Electric Fans (EFs) was proposed. The aircraft was sized for an interval design ranges 900– 2100 nm and the resulting performance was compared against a projected, solely combustion-based aircraft. The analysis in terms of relative change in block fuel reduction and change in vehicular

21

efficiency indicated a significant block fuel reduction could be achieved with increasing HP,USE. However, the potential block fuel reduction vanishes for design ranges above 1700 nm assuming a battery gravimetric specific energy of 1.5 kWh/kg. The installation weight of the batteries and the electrical system result in a large weight increase of the aircraft. This effect, which increases with the level of HP,USE, counteracts the benefit of an increase in overall propulsion system efficiency afforded by the utilization of the more efficient electrical system approach. As a consequence, the vehicular efficiency of the hybrid-electric aircraft degrades rapidly with increasing HP,USE and was found to be lower than the projected, solely combustion-based aircraft for all investigated design ranges. The sensitivity of the design with regards to the battery gravimetric specific energy was also assessed. Intuitively, for lower battery gravimetric specific energies, the amount of batteries increases for a given energy demand, which penalizes strongly the performance with regards to fuel and energy consumption as well as aircraft weight. Assuming a battery gravimetric specific energy of 1.0 kWh/kg, no real potential in block fuel reduction was displayed for design ranges above 1100 nm. Moreover, the narrowbody fuselage shape imposes a volumetric constraint for the housing of the batteries. Evolution of the fuselage geometry towards doublebubble cross-section might be required for required battery housing as well as to provide similar cargo volume standards afforded by contemporary state-of-the art aircraft. The influence of the HP,USE to flight technique optimization was also analyzed in this technical article. A reduction in optimum flight-speed and altitude was assessed with increasing HP,USE. However, by mimicking the Long Range Cruise definition strategy, these effects would not affect strongly the flight technique optimality of hybrid-electric aircraft compared to conventional gas-turbine only aircraft. The operational scheme of the hybrid-electric system employed in this concept could be revisited in the way that the GTFs would be optimized and would operate at its peak efficiency during cruise, whereas the fluctuation of thrust required would be alleviated by the EFs. In this way, the GTF efficiency would not be impaired by any variations in HP,USE. As the efficiency of an High Temperature Superconducting motor remains almost constant for broad range of power levels, the thrust throttling of the EFs would not impair the overall propulsion system efficiency. The sizing cascading effect will depend on the sizing of the electrical system and the total battery mass required. The investigation of this operational strategy will be subject of future work. In order to compensate the negative weight impact of the advanced electrical system integration, more tightly-coupled approaches need to be conducted in further work. By enhancing the synergies between the propulsion system, aerodynamics and the structure, the field of distributed propulsion will definitely play a major role in establishing the feasibility of hybrid-electric technologies. Furthermore, innovative combinations between energypower generation and the fuel-battery hybrid-electric system needs to be further assessed. Moreover, a deeper investigation of possible synergies provided by the electrical propulsion system integration on another systems on board of the aircraft need to be considered.

Acknowledgments The authors would like to thank Sascha Kaiser for fruitful discussions and valuable advice.

Appendix A. Supplementary material Supplementary mkaterial associated with this article can be found in the online version at http://dx.doi.org/10.1016/j.paerosci.

Please cite this article as: C. Pornet, A.T. Isikveren, Conceptual design of hybrid-electric transport aircraft, Progress in Aerospace Sciences (2015), http://dx.doi.org/10.1016/j.paerosci.2015.09.002i

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2015.09.002. [33]

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Please cite this article as: C. Pornet, A.T. Isikveren, Conceptual design of hybrid-electric transport aircraft, Progress in Aerospace Sciences (2015), http://dx.doi.org/10.1016/j.paerosci.2015.09.002i