PROGRESS AND PERSPECTIVES OF ELECTRIC AIR TRANSPORT · 2018-06-22 · Progress and Perspectives of...

PROGRESS AND PERSPECTIVES OF ELECTRIC AIRTRANSPORT

H. Kuhn, A. Seitz, L. Lorenz, A.T. Isikveren, A. SizmannBauhaus Luftfahrt e.V., Lyonel-Feininger-Str. 28, 80807 Munich, Germany

[email protected]; [email protected];[email protected]; [email protected];

[email protected]

Keywords: electric motive power systems, Ragone metrics, exergy concept, electric flying

Abstract

We address the expediency and explore the inno-vation potential of Hybrid Electric Power Archi-tectures (HEPAs) in aviation. As a starting point,we define three fundamental principles for the fea-sibility assessment of electric flight and introducethe Ragone metrics as essential analytical param-eters. Following a discussion of key system com-ponents and their technology perspectives, we de-scribe four examples of HEPA systems, includingthe Universally-Electric System Architecture (U-ESA), which realizes zero in-flight emissions. Thesignificant loss in aircraft range brought about bythe high weight of U-ESA components can –inview of present technology options– only be coun-teracted by providing part of the aircraft’s powerdemand with a conventional turbo-engine. Weanalyse this ‘battery + fuel’ hybrid system in detailand draw interesting conclusions about the impactof the degree of hybridization on aircraft range, en-ergy consumption and CO2 emissions.

1 Introduction

Future growth of the aviation industry will de-pend crucially on the reduction of its environmen-tal footprint through substantial efficiency gainsand substitution of fossil kerosene by renewablealternatives [1–3]. The universally electric aircraftis the most radical step in this direction: If battery-powered, such an aircraft would have zero in-flight

emissions, surpassing even the ambitious Flight-path 2050 [4] goals of 75% reduction in CO2 and90% less NOx. The flexibility in choosing the pri-mary source of electric energy paves the road tolong-term sustainability and a low overall emissionbudget through the use of power from solar radia-tion, wind, water or geo-thermal sources.

The electric aircraft concept also offers thepossibility to optimize aircraft efficiency using rad-ically different configurations, due to the new de-grees of freedom released by the separation ofpower generation from power consumption. Aero-dynamic as well as propulsive efficiency can bene-fit from the liberty of integration offered by univer-sally or hybrid electric power systems, as can con-trol dynamics, service accessibility, and mass andload distribution. Together with the future tech-nology perspectives of electric components andsub-systems, these advances in configuration andpropulsion integration can potentially yield thestep changes in efficiency and emission necessaryto ensure sustainable long-term growth for the avi-ation sector.

In this paper, we first address the fundamentalfeasibility and scaling properties of electric flight,highlighting the importance of properly definedsystem boundaries and introducing the Ragone di-agram as the key tool for system comparison. Wethen take stock of the various technologies re-quired to set up Hybrid Electric Power Architec-tures (HEPA) and discuss their potential for futuredevelopment. We illustrate four distinct system ar-

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H. KUHN, A. SEITZ, L. LORENZ, A.T. ISIKVEREN & A. SIZMANN

chitectures: an entirely battery-powered system, afuel cell system and their hybrid combination, i.e.three options for a Universally-Electric System Ar-chitecture (U-ESA); and a hybrid of a conventionalturbo-engine and battery. The performance of a U-ESA is then analyzed in the context of concrete fu-ture mission requirements [5]. Finally, we presenta detailed study of the possible range extensionsfor an U-ESA platform in hybrid combination witha conventional turbo-engine: Introducing the De-gree of Hybridization (DoH) as a key parameter,we show how energy consumption and CO2 emis-sions per passenger-kilometer vary with DoH, con-cluding that a given target for emission reduction–such as set by the EU in the Flightpath 2050 re-port [4]– essentially dictates a minimum amount ofhybridization for the power system.

2 Feasibility and Scaling Properties of Elec-tric Flight

A major challenge for electric flight lies at the in-terface between scientific principles and develop-ments related to electro-mobility outside the fieldof aviation, and aviation-specific questions such as(novel) aircraft configurations and optimal systemintegration. Only a successful combination of bothaspects will allow to exploit their full potential forinnovation. Here, we set the stage by introducingthe underlying guidelines and parameters for oursubsequent analysis.

2.1 Principles of the Feasibility Assessment ofElectric Flight

Three fundamental principles of the feasibility as-sessment of electric flight as proposed by BauhausLuftfahrt are [2, 3]:

1. Exergy concept: The usefulness of an en-ergy carrier for aviation is determined by itsspecific exergy content, i.e. gravimetric andvolumetric exergy density, rather than by itsenergy content.

2. Specific power and Ragone metrics: Theusefulness of an energy carrier in combina-tion with a power converter is fundamentally

determined by their combined power densityand exergy density (Ragone metrics). Thesetwo metrics are the key indicators for electricaircraft feasibility when comparing alterna-tive power sources.

3. Hybridization degrees of freedom: Two ormore energy storage and power conversiondevices that separately are inadequate forelectric flight (according to their Ragonemetrics) may constitute an enabling powersystem when combined into a hybrid system.

Each subsequent principle adds to the previous onean essential and new dimension in the feasibilityassessment of electric flight. We now discuss thesignificance of these principles in detail.

The first principle is based on the second lawof thermodynamics and relates the fact that ex-ergy is the part of the energy that can absolutelybe converted into useful work. The net exergyof a given system is determined by the sequenceof conversion efficiencies in each of its individualcomponents. Note that, unlike universally elec-tric power systems, combustion engines are funda-mentally limited by the Carnot efficiency. There-fore, the huge apparent energy density gap be-tween kerosene and state-of-the-art batteries forelectromobility (a factor of 50 to 60) is reducedto a factor of 20-30 when one accounts for the ab-sence of the Carnot limitation in non-thermal elec-tric power conversion.

According to the second principle, the feasibil-ity of electric flying is best understood in terms ofthe Ragone diagram as shown in Fig. 1. A highenergy or exergy density, such as it is associated,e.g., with hydrogen-powered fuel cell systems, isan insufficient (and even misleading) indicator forthe capability to realize electric flight because thespecific power requirements of take-off and climbhave to be met at the same time.

These requirements are increasingly difficult tofulfill for larger aircraft take-off masses. In somecases the limitations of fuel cell systems are bestovercome by including an auxiliary high-power-density battery system. Hence, a hybrid systemas advocated by the third principle can meet thepower and energy requirements expressed in terms

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Progress and Perspectives of Electric Air Transport

of the Ragone metrics even though the indivdualsubsystems show inadequate performance charac-teristics.

2.2 System Boundaries and Ragone Represen-tation

The exergy densities and power densities of twoexemplary motive power systems are displayed inthe Ragone diagram of Fig. 1. ‘Battery+Motor’denotes a universally battery powered system withelectric motor (dashed light blue line), while‘FC+BoP+H2+Tank+Motor’ refers to a fuel cellsystem (BoP: Balance of Plant) powering an elec-tric motor (using a liquid hydrogen tank as anenergy source, solid green line). Also depictedis the Ragone representation of typical medium-range aircraft with conventional turbo-engine: Thepoints A,B,C (orange triangles) correspond to se-lected points on the aircraft’s payload-range dia-gram (compare inset plot in Fig. 1). During a typ-ical mission, the aircraft will follow the solid or-ange line starting in point A. Note that A and Bare very close, and that the aircraft ‘crosses’ theU-ESA (‘Battery+Motor’) performance line whilepassing from point B to the state where only re-serve fuel is left in the tank (point C).

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101

102

103

104

10510

0

101

102

103

104

105

FC+BoP+H2+Tank+Motor

Specific Exergy at Power System Level / Wh/kg

Spe

cific

Pow

er a

t Pow

er S

yste

m L

evel

/ W

/kg

Battery+Motor Turbo−FanFuture

PropulsionSystem

A

B

C

Range

Pay

load

A

BC

Fig. 1 Ragone diagram presentation (specific ex-ergy [Wh/kg] vs. specific power [W/kg]) of differ-ent HEPA systems and the ‘Ragone range’ of thepayload-range diagram of a future medium rangeaircraft (see inset plot). Data from Table 1.

Table 1 Characteristic values for the power sys-tem components.

Component Unit 2010 2035+

Kerosene kWh/kg 11.9 a

Hydrogen kWh/kg 33.31 a

Fuel tank kg – b

Lithium battery kWh/kg 0.60 1.5Turbo-enginec kW/kg 15.0 18.0Fuel cell (PEFC) d kW/kg 1.2 3.0Percentage of BoP e 0.40 0.30Generator kW/kg 10.0 20.0Motor kW/kg 10.0 20.0Power managementf kW/kg 10.0 15.0

Efficiencies 2010 2035+

Turbo-engine 0.50 0.55Fuel cell 0.50 0.65Generator 0.99 0.99Motor 0.99 0.99Power management 0.95 0.98a Lower Heating Value (LHV)b The mass of the fuel tanks are calculated separately

depending on the fuel amount.c bare turbo-shaft specific power at max powerd PEFC: Polymer Electrolyte Fuel Celle BoP: Balance of Plantf single system without redundancies

2.3 Architectures and Combinations

The power systems plotted in Fig. 1 are only twoamong a variety of architectures that are possi-ble for HEPA systems. In principle, the HEPAconcept includes all system combinations that canbe formed with the various sources of energy andtypes of power converters in Fig. 2. ‘PMAD’ inFig. 2 stands for Power Management and Distribu-tion System, i.e. the components that establish theessential link between power generation and con-sumption in the turbine. The performance char-acteristics for each system are listed in Table 1.These numbers are the basis for the calculationspresented in the later sections of this paper.

A useful exergy analysis of these systems re-quires a proper identification system boundariesand estimate of conversion efficiencies. Further-more, the scaling behaviour and feasibility assess-

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ment requires the Breguet and the Specific BreguetRange equations to be adjusted to electric batterypower release without change in mass. These as-pects have been carefully considered when estab-lishing the systems’ representation in the Ragonediagram of Fig. 1.

3 Future Perspectives of Key Technologies

For an accurate evaluation of the feasibility,progress and perspectives of universally elec-tric and hybrid motive power systems a physics-based fundamental understanding and modelingapproach of the basic components is mandatory.We therefore examine the prospects of key tech-nologies relevant for electric flight – such as bat-teries, fuel cells, motors and generators –, pointingout selected scientific and technical limits on theirfuture development.

3.1 Batteries

Today, standard electrode materials such as lithiummetal oxides LiMO2 (M = Co, Mn, FePO) andgraphite (C) limit the specific energy of commer-cial batteries to a level below 300 Wh/kg. Never-theless, existing batteries have been demonstratedto serve as the main source in small light airplanesenabling flight times of up to 3 hours.

Crucial battery components with significant in-novation potential are the electrodes and the elec-trolyte. Therefore a range of new electrode ma-terials is under investigation promising significanthigher capacities as well as charge-discharge rates.As stated elsewhere, the specific power is as im-portant as the specific energy in order to providehigh power capabilities at high energy storage lev-els to allow for electric flight [2].

Recent developments and laboratory-scaledemonstrations in electrode materials provepromising performance enhancements of sec-ondary batteries that could result in a significantstep change regarding energy and power densities.Kang et al. demonstrated very high discharge ratesfor lithium phosphate coated LiFePO4 electrodesallowing for a full charge between <2 and 30minutes at reasonable specific capacities of 130

to 170 mAh/g, respectively [6]. LiNiMnCo oxideelectrodes are another promising class of positiveelectrodes. Ban et al. demonstrated charge-discharge cycles at 5 C (full charge in 12 minutes)at reasonable capacities of 120 mAh/g after 500cycles, whereas at 10 C (6 minutes) still capacitiesof 110 mAh/g after 500 cycles are feasible [7]. Forcomparison, standard LiCoO2 electrodes exhibita capacity of less than 135 mAh/g at rates below1 C (1 hour). Sulphur has a theoretical capacityof 1166 mAh/g which is far beyond the capac-ities of the previous discussed materials and istherefore a promising material. The drawbacks ofsulphur are its insulation characteristics and largevolume changes upon lithium insertion/extraction.Electric conducting additives such as carbonneed to be added, hence reducing the specificcapacity. The volume changes may be overcomeby nano-structuring of the electrode. Yang et al.realised a cell incorporating a lithium-sulphurcarbon composite electrode having a specificenergy of 315 Wh/kg at battery level [8].

To enhance the performance capabilities of abattery both electrodes need to be optimised interms of specific charge and rate capability. Com-pared to the positive electrode there are more ma-terials under investigation for the negative elec-trode such as silicon as the most popular today dueto the theoretical specific capacity of 4200 mAh/gand significant improvements in nano-structuring[9–13]. Chan et al. demonstrated capacities be-tween 2100 and 3500 mAh/g for silicon nanowiresat discharge rates of 1 C to 1/5 C, respectively. Forcomparison standard graphite electrodes exhibit acapacity of 350 mAh/g at similar charge-dischargerates.

In the past, the specific energies of batteriesincreased by 7 % per year by optimising standardelectrode materials and their structure. Accordingto the maximum theoretical capacities of the elec-trode materials this trend will not continue due tophysico-chemical limits of the materials. We set upan electrode inventory comprising 9 positive and20 negative electrode materials and investigated allpossible combinations of these electrodes in or-der to identify their potentials of specific energywith respect to equilibrium potential, molar mass

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Fig. 2 Components for power generation (left) and power consumption (right) from which a combina-torial variety of electric power architectures can be derived.

and reversible range [14]. Specific energies in therange of 1000 to 1500 Wh/kg seem to be feasiblein the mid to long term time horizon which is 5 to 7times the specific energies of commercial batteriestoday.

3.2 Fuel Cells

Polymer electrolyte fuel cells (PEFC) are con-sidered as possible power generators because ofthe advantage of using high energy density liq-uid hydrogen in combination with a high fuel-to-electricity conversion efficiency. PEFCs providea specific power of up to 1.25 kW/kg (at stacklevel), the highest among all fuel cell technologies.A study by Middelman et al. shows that specificpowers of 2 kW/kg at stack level are feasible in thenear term [15] by improving bipolar plate charac-teristics, the main contributor to weight and vol-ume of the stack.

Platinum is still the main catalyst for fuel cellelectrodes. The platinum loading greatly reducedfrom initially 4 mg/cm2 to 0.3 mg/cm2 retainingsame or enhanced power capabilities at signifi-cantly reduced costs. Regarding the CO intoler-ance of platinum it is of major importance to re-move any traces from the operating gases. Newcatalysts such as platinum supported on a TiWOsubstrate exhibit higher tolerance levels towardsCO [16]. Core-shell Pd/FePt nanoparticles serve asalternative catalysts for the oxygen reduction reac-tion (ORR) [17]. These new catalysts materials al-low for enhancing the durability, stability and costof PEFCs.

Nafion is the state-of-the-art membrane forPEFC. These membranes need to be humidifiedin order to conduct hydrogen ions and thereforethe humidity management of a PEFC is crucial forits power capabilities. Based on alternative sili-con membranes the humidity management may be

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significantly simplified, hence reducing balance ofplant requirements [18]. Increased ionic conduc-tivity further reduces internal losses and hence in-creases specific power of the PEFC stack.

Considering improvements in catalyst loading(less catalyst per area), higher ionic conductiv-ity based on alternative membranes (less inter-nal losses) and enhanced bipolar plates the spe-cific power of PEFC may be in the range of 2 to3 kW/kg in the future.

3.3 Electric Motors

The third key component is the electric motor pro-ducing shaft power to drive the propulsor. Normal-conducting motors have specific powers in therange of 2 to 10 kW/kg, entering the power densitydomain of turbo-engines.

Power density prospects of high-temperaturesuperconducting (HTS) motors as high as 40kW/kg, i.e. well above the level of state-of-the-art turbo-engines, have been reported [19]. Aphysics-based model, extended to the supercon-ducting regime, will provide a future-proof frame-work for power and mass scaling of HTS electricmotors, and analogously of generators for turbo-electric systems as well.

4 Electric Power System Architectures

Among the choices illustrated in Fig. 2, we nowdiscuss four power system examples in moredetail, highlighting their advantages and chal-lenges. The universally-electric battery system(Section 4.1) and its hybrid with a conventionalturbo-engine (Section 4.4) will be studied in a con-crete integration example in the next two sections.

4.1 Battery System

A typical battery-powered U-ESA is shown in Fig.3; note that energy storage and conversion arecombined inside the battery, making it the sys-tem with the least number of components. Twodifferent types of batteries, which can be used in(parallel) combination, are indicated, namely highenergy density-optimized batteries (‘E’) and their

high power density counterpart (‘P’). Unlike con-ventional turbo-engine aircraft, an airborne inte-gration platform with batteries does not changeits weight in flight (since no fuel is burned off),which will have a significant impact on structureand aerodynamics (see Section 6.2). The main

Fig. 3 U-ESA with batteries as only energy source.

performance penalty for purely electric power sys-tems at the moment and in the foreseeable futureis the high battery weight (see also Table 2). Tech-nology advances in other mobility sectors an helpto bridge the gap between the conflicting require-ments of high energy and high power density, butstep changes are necessary to make electric flight areality beyond the scale of single-seater airplanes.

Table 2 Characteristics of battery-powered U-ESA.

Positive Negative Interesting

+ very high effi-ciency (∼90%)

- high batteryweight

◦ recycling of bat-teries

+ no CGa move-ment

- landing weightsame as take-offweight

◦ zero globalemission possible

+ low mainte-nance

- small range ◦ replacement /recharging duringturn-around

+ zero emissionsin flighta Center of Gravity

4.2 Fuel Cell System

Figure 4 depicts the fuel cell power system; withits linear path from the hydrogen tank through themotor to the propulsor, this system is very simi-lar to the conventional aircraft power plant, withkerosene replaced by hydrogen and the fuel cell in-stead of the turbo-engine. However, the efficiency

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of energy conversion is significantly higher for fuelcells. Note that in this system, there is mass changeduring flight due to fuel burn; however, because ofthe very low gravimetric density of hydrogen, en-ergy storage contributes to the overall aircraft massbudget.

Fig. 4 U-ESA with fuel cell and hydrogen tank.

Table 3 Characteristics of fuel cell-powered U-ESA.


+ higher effi-ciency

- H2 tank integra-tion

◦ zero CO2 possi-ble

+ peak power fex-ibility of fuel cells

- new fuel infras-tructure needed

◦ on-board use ofproduced water

4.3 Fuel Cell Battery Hybrid System

This hybrid system (see Fig. 5) combines the in-trinsic advantages of the battery motor and the fuelcell system architectures. Notably, this system willprofit from technology advances in both fields ofresearch. Interesting synergies can be achieved by,for example, using the batteries as an emergencypower source in case of fuel cell failure or by thefact that the cooling architecture of the hydrogenstorage may be beneficially extended to supercon-ducting PMAD components.

Fig. 5 U-ESA with hybrid hydrogen-supplied fuelcell and batteries.

4.4 Battery Turbo-Engine Hybrid System

In view of the analysis to be performed in Sec-tion 6, as a final example we discuss a hybrid

of a battery-powered U-ESA with a conventionalturbo-engine. The corresponding system is shownin Fig. 6. While any system using conventionalcombustion for power generation must miss theultimate goal of zero in-flight emissions, this hy-brid solution offers a feasible solution to the rangeloss problem caused by high U-ESA componentweights. In Section 6, where notably we will intro-duce the Degree of Hybridization (DoH) to charac-terize the amount of power provided through eachpower conversion path.

Fig. 6 Parallel hybrid system architecture with bat-tery and conventional turbo-engine.

Table 4 Characteristics of battery turbo-engine par-allel hybrid system.


+ higher ef-ficiency thanconventionalturbo-engine

- weight penaltyof PMAD, motorand batteries

◦ integration-dependent struc-tural benefits

+ fuel system andinfrastructure par-tially unchanged

◦ noise reduction

5 Requirements for Future Electric Aircraft

Transport aircraft employing electro-mobility as ameans of motive power have the benefit of realiz-ing zero-emissions for both ground maneuveringand in-flight operations. Notwithstanding this de-sirable outcome, the detrimental effect of dramaticrange reduction compared to the operational reachafforded by gas-turbine propulsion systems to-day makes the engineering systems trade-off pro-hibitive. The excessive weights of electro-mobilitybased solutions are attributable to high specific

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Fig. 7 The CeLiner platform upon which all U-ESA sub-systems were designed.

weights of components that describe the integratedsystem, i.e. energy source (whether battery-only,fuel-cell-only, or a combination of the two), con-trollers, wiring and associated buses/connectors,and motors. Unfortunately, studies show thatthis penalizing circumstance cannot be effectivelytraded away by the rather high component efficien-cies inherent to them. In this context, since weightis at a premium, a means of off-setting the detri-mental effects of high take-off/landing regulatedweights and en route flight gross weights wouldbe to maximize low-speed and high-speed lift-to-drag ratios as well as find a means of improvingthe structural weight efficiency of the aircraft.

5.1 Mission Requirements

In order to perform the requisite array of sub-systems engineering trade-studies, the CeLinerwas procured [5], which allowed to address siz-ing sensitivities and establish robustness. The aimwas to understand the potential for scalability ofthe required technologies and at the same time en-courage their propensity for evolutionary develop-ment. For purposes of showing characteristics as-sociated with electro-mobile transport aircraft so-lutions, a pre-concept design of a 190-passengeraircraft aimed primarily at the short-haul marketwas synthesized (see Fig. 7). Exceeding Flight-path 2050 goals [4], the unique feature of this ini-tial technical assessment activity was to produce acommercial transport that delivers zero-emissionsfrom gate to gate. Entry-into-Service (EIS) for thebaseline was set as year 2035.

This ambitious zero-emissions target wasachieved through implementation of a Universally-Electric Systems Architecture (U-ESA), which

includes a propulsion solution solely poweredthrough electrical means. Advanced Li-ion batter-ies constitute the only source of electrical power.The batteries have been configured as modularpacks that fit into suitably modified LD-3 con-tainers in order to facilitate expedient turn-aroundand aid loadability. With regards to motive powerthe design proposal utilizes ducted fans run byHigh Temperature Superconducting (HTS) electricmotors reflecting posited performance based upontest-rig data obtained from the literature [20, 21].

An innovative C-Wing morphology, dispens-ing with the need for a dedicated horizontaltailplane as well as maximizing vehicular aerody-namic efficiency, was selected due to airport com-patibility requirements that imposed a span limit.Also, for purposes of ensuring compliance withcontemporary standards of Configuration, Main-tenance and Procedures (CMP), a complete air-craft systems design adhering to 90min ExtendedTwin OPerationS (ETOPS) was stipulated. Thislast technical requirement infers the U-ESA sys-tem needs to be supplied by at least three indepen-dent sources of power.

5.2 Universally-Electric System Require-ments

The contemporary approach to electrical systemsintegration, the so-called ‘more-electric’ and the‘all-electric’ architectures, irrespective of the ex-tent of electrification require fundamentally me-chanical off-take from a gas-turbine power source.Over time, incremental improvements will beachieved through wholesale application of DirectCurrent (DC) only power transmission. Apartfrom the principal advantage of improving powerloss reduction through increased voltages, DC sys-tems minimize electro-magnetic interference andare compatible with battery and/or fuel-cell energysources since no specialized (and heavy) equip-ment for conversion to-and-from Alternating Cur-rent are required.

One engineering proposal for future ubiquitouselectrical systems architectures is the U-ESA. TheU-ESA is divided into three different main sys-tems, all representing different power and voltage

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sources. The first level, dedicated to the propul-sion system, is characterized by high power andvoltage. The second level caters to onboard cus-tomer system requirements: i) non-essential, e.g.galleys; ii) essential, e.g. flight control; and iii) vi-tal, e.g. emergency lights. This is associated witha medium power and voltage. The third level, withrelatively low power and voltage demand feeds theavionics equipment.

6 Technology Gaps and Future Electric AirTransport Perspectives

It has been stated earlier that propulsion basedupon the precepts of electro-mobility suffer fromvehicular efficiency degradation due to relativelyhigh specific weights of U-ESA-type components,thereby producing less than desirable operationalreach, namely, range. It was deemed appropriateto perform a design perturbation study with in-tent to show the relative merits of implementingpropulsion system solutions comprising a varyingmix of gas-turbine and HTS motor for a requisitethrust. The following section offers details aboutthe problem formulation behind this design pertur-bation and sensitivity study, and more importantly,offers a number of salient conclusions that can bedrawn from the results.

6.1 Hybridization Degree of Freedom

In order to identify degrees of freedom for extent-ing the limited range of electrically powered air-craft, a hybridization study was performed. To thisend, the universally-electric, short-range, medium-capacity aircraft concept CeLiner was hybridizedthrough additional conventional energy storageand conversion, i.e. kerosene fuel and turbo-shaftgas-turbines partially delivering the required en-ergy and power for the aircraft (compare Fig. 6).

CeLiner reflects technology status 2030 interms of aerodynamic and structural performance.The energy and propulsion system of the refer-ence aircraft comprises two installed power plants,an ETOPS compatible PMAD and battery system.The installed power plants consist of the bare HTSelectric motors, the required cryocoolers and mo-

tor controllers, as well as the propulsive devicesincluding the fan components, the fan drive gearboxes and the nacelles. The maximum take-offweight (MTOW) share of the overall energy andpropulsion system is 42.8%, while the maximumstructural payload is 19.0% of MTOW.

For the hybridization study, MTOW, struc-tural weights and aerodynamic properties werekept fixed, equal to the reference aircraft, thereby,neglecting potential structural synergies or draw-backs due to the varying overall system architec-ture. With the assumption made, the overall in-stalled power was considered invariant for the dif-ferent aircraft investigated.

For the sizing of the individual components ofthe energy and propulsion system the individuallyrequired maximum powers were considered as sig-nificant. The corresponding critical sizing casesare indicated in Table 5.

Table 5 Summary of component modelling for hy-brid energy and propulsion system.

Sizing Case Effi-ciency

PowerDensity

EnergyDensity

Batterya take-off, AEOc const const constPMADa take-off, AEOc const const n/aHTSmotorb

take-off, OEId const f(Pel) n/a

Gas-turbine

take-off, OEId f(Pco) f(Pco) n/a

Fuel n/a const n/a consta incl. redundancies for ETOPSb incl. cryocooler and controllerc All Engines Operationald One Engine Inoperative

In the study, the degree of hybridization, H ,was defined, expressing the share of installedpower between electrical, Pel, and conventional,Pco, systems:

H =P − Pel

Pco − Pel

, where P = Pco + Pel . (1)

All components of the energy and propulsion sys-tem were sized according to the correspondingshare of overall required power. The componentefficiencies and power densities, i.e. power percomponent mass, were expressed as functions of

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the individual sizing powers. The specific powercharacteristics versus installed power for HTS mo-tors were taken from Masson et al. [20, 21] andextrapolated to technology 2030+. The specificpower of battery and PMAD system was treatedconstant for variations in power sizing. Using thedetermined power densities, the masses of the en-ergy transmission components, i.e. PMAD, elec-tric motors and gas turbines, were determined. Thesizing power of the battery system was treated asdirectly coupled to the PMAD system, i.e. massand available energy from the battery system scaleproportionally with PMAD sizing power. Theresidual mass budget for the overall energy andpropulsion system was considered available as hy-brid energy storage, i.e. kerosene fuel and tur-boshaft engines adding to power of driving thepropulsive devices. The masses and efficiencies ofthe propulsive devices were considered unaffecteddue to invariant aircraft thrust requirements. Theefficiencies of the electrical system componentswere assumed independent from sizing power vari-ations, in the first instance.

The dependency of turbo-shaft engine powerdensity on installed power was mapped based onan empirical correlation given in Ref. [22]. Thecorrelation was calibrated to match advanced en-gine characteristics in the relevant shaft powerclass. The power density of fuel was consid-ered unlimited while potential tank and fuel sys-tem masses were accounted for within the chosenvalue of fuel energy density.

Aircraft range was calculated using the Bre-quet equation, extended for hybrid energy, basedon the available exergies of the energy storage me-dia and the corresponding individual mass lossesduring flight mission.

The integrated results of the study are shownin Figure 6.1, including aircraft range and storedexergy, as well as, CO2 emissions and energy con-sumption per passenger and 100 km. This figureoffers some interesting insights when it concernsselection of DoH in the context of passenger trans-port propulsion integration. At first glance it canbe observed that the rate reduction in range capa-bility for increasing level of electrification (repre-sented by the DoH values tending towards zero) is

non-linear tending towards vanishing slope.As a primary effect, the rate reduction in range

is directly proportional to the reduction in to-tal available exergy of the turbo-electric system,since the available exergy from kerosene is an or-der of magnitude larger than what can be drawnfrom a battery based energy source. Other aspectsthat come into consideration are effects relatedto higher Instantaneous Gross Weights (IGWs),which is a bi-product of increasingly electrifiedpropulsion system installations, i.e. PMAD ver-sus fuel system, and the lower power density ofinstalled HTS motors vs. advanced gas-turbines.

For turbo-electric power mixes of H ≤ 0.35(65% HTS motor and 35% gas-turbine for installedthrust), the rate reduction in range tends to dimin-ish noticeably, i.e. less range loss with increasingamount of electrification. This can be explained bythe interaction of steadily increasing hybrid systemefficiency (diminishing gas-turbine efficiency dueto scale effects versus a constant HTS motor effi-ciency), reduction of the penalizing impact of highIGW (PMAD weight increase offset by steadilyimproving HTS motor power density), and, in-creasing available exergy from the batteries.

Regarding the CO2 per PAX·100 km emissionstrend with diminishing DoH, one can observe thatparity occurs for turbo-electric hybrids of around50% (H = 0.50) with that of the all-gas-turbine(H = 1.0). A 50% mix of turbo-electric propul-sion yields range capability of around 3000 nm.For typical narrow-body aircraft this constitutes avery reasonable level of operational reach. It ishighlighted that the all-gas-turbine solution cor-responds to an evolutionary internal combustionpropulsion system that already delivers approxi-mately 25% improvement over the year 2000 state-of-the-art. The weak impact of DoH until the 50%mix point indicates that propulsion systems biasedtowards all-gas-turbine installations have no dis-cernible benefits in the context of emissions - onlyyielding a detriment in terms of range capability.

If one now considers a CO2 per PAX·100km emissions target in-line with propulsion-onlyexpectations according to Flightpath 2050 [4],namely, a total reduction of 40% compared to theyear 2000 datum [23], this corresponds to around

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0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10

1000

2000

3000

4000

5000

6000

7000

Degree of Hybridisation H = (P-Pel)/(Pco-Pel) [-]

Airc

raft

Ran

ge [N

M]

Reference A/C:Full-Electric

Design Range 900NMMTOW 109.3t

Payload 190Pax0

2

4

6

8

10

Rel

. Exe

rgy

Sto

rage

Cap

acity

Ex re

l=(E

x el+E

x co)/E

x el [-

]

0.0

2.0

4.0

6.0

8.0

10

CO

2 Em

issi

ons

per P

AX

per

100

Kilo

met

er [k

g/P

AX

/100

km]

10

15

20

25

30

Ene

rgy

Con

sum

ptio

n pe

r PA

X p

er 1

00km

[kW

h/10

0PA

Xkm

]

Rel. Exergy StorageAircraft Range

CO2 per PAX

Energy per PAX

Fig. 8 Degree of hybridization for a parallel hybrid of gas-turbine and electric system; H = 1.0 denotesall-gas-turbine and H = 0 denotes universally-electric propulsion.

H = 0.25 (75% HTS motors and 25% gas-turbinefor requisite thrust) and range capability of approx-imately 1750 nm. Beyond this, if it is assumedthe propulsion system alone should deliver Flight-path 2050 goals, i.e. a 75% reduction in CO2 perPAX·100 km, this corresponds to an operationalreach of just over 1000 nm. Finally, based on thepresented study, a full electro-mobility solution,that is to say zero-emissions in flight, would pro-duce a range capability of 900 nm.

In summary, it can be concluded that H =0.25 would meet the propulsion-only targets set byFlightpath 2050 (40% reduction compared to year2000 datum) and still produce a reasonable max-imum range (1750 nm) for the category of trans-port aircraft presented in this investigation. It isour view that if the ambition is to realize an evengreater reduction in CO2 per PAX·100 km, thetrade-off between hybrid system complexity andrange potential is not sufficiently in favor of hybridturbo-electric solutions, and thus an all-electricpropulsion system design should be considered ifimproved architectural solutions for hybrid energyand propulsion systems are not available.

6.2 Impact of Structural and AerodynamicEfficiency Improvements

A prediction algorithm suited for the pre-designstage of transport aircraft design sizing was appliedto the DoH study in order to gauge the relative mer-its of improved structural and/or aerodynamic effi-ciencies. The prediction method is a combinationof statistical correlations of design variables andmacro-objective functions with fractional changeanalytical constructs [24]. The analytical compo-nent of the fractional change method operates withthe underlying premise that the designer/analystbegins with a seed condition or aircraft. Briefly,by considering an increment in variable x as dx or∆x, a fractional change to a new value, x1, smallor otherwise, in a seed parameter x0 is defined as

/x =∆x

x0

=x1 − x0

x0

(2)

Now, derivation of an integrated range model be-gins with the Breguet equation for range (R). Indifferential form the Breguet equation accounts forthe change in aircraft instantaneous gross weight,W , based on an instantaneous specific air-range.One variation of the differential equation for rangeis to express it as a function of fuel calorific value,Hc, the overall power plant efficiency, η and the

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vehicular lift-to-drag ratio, (L/D), i.e.

dR =Hc

gη · (L/D)

dWfuel

W. (3)

Eq. 3 quotes only increments in range for an in-stantaneous condition. The actual aim is to seekan integrated range result wherein the product ofη · (L/D) continually varies for the entire mission.One method is to assume an arbitrary referencepoint such as Initial Cruise Altitude (ICA) is re-lated to the Takeoff Gross Weight (TOGW), whereW = kclb · TOGW , kclb < 1, and incorporate aprofile correction (Θprf) that captures step-cruisingas well as continually varying η ·(L/D) effect. Theprofile correction coefficient value will be uniquedepending on the relationship between W at theICA and the IGW at end of mission. In view ofthis association and since account is made for Wat initial cruise in Eq. 3, it is reasonable to con-clude that proportionality will exist between Θprf

and Zero-Fuel Weight, i.e. ‘ZFW = Basic Oper-ating Weight (BOW) plus payload’, and a suitablerelationship has been established to be an exponen-tial one, i.e. Θprf = (1+/ZFW )ε. In order to per-mit an even wider scope of functional sensitivitywith sufficient accuracy, an extension to the pro-file correction [24] would also include account ofvariations due to change in useful load as payloadremains fixed, i.e. TOGW varies. This relationshipwas also found to be suitably represented by an ex-ponential one with Θprf = (1 + /ζflt)

α, where /ζflt

represents the flight fuel fraction, or,

/ζflt =(1 + /Wflt)

(1 + /TOGW )− 1 . (4)

Upon introduction of an extended Θprf definitioninto the integration of Eq. 4, the method to quantifyfractional change in range becomes

/R =(1 + /η) [1 + /(L/D)] (1 + /ζflt)

α

(1 + /ZFW )ε− 1 .

(5)Eq. 5 permits opportunity to perform sensitivitystudies such that one can inspect the merits of in-creased structural efficiency (reduction in the ZFWvariable) and higher aerodynamic efficiency (in-crease in (L/D) parameter). It should be noted

that for an all-battery electro-mobility systems theexponents α and ε would be taken to be unity sincefor the IGW at any point during the flight is con-stant, i.e. equal to TOGW.

Figure 9 displays the functional sensitivity ofR associated with arbitrary variations in ZFW (orTOGW if H = 0 and an all-battery energy sourceis utilised) and (L/D). The chart is applicable forall candidate DoHs presented in Fig. 8. If, for in-stance, a H = 0.25 is assumed (recall discussionof Fig. 8 in Section 6.1), and, an aerodynamic effi-ciency of +10% together with a reduction in ZFWof -5.0% is considered, the range capability of thiscandidate propulsion system mix would increaseby 17.8%. In other words, the range would in-crease by around 300 nm from the 1750 nm datum.

7 Conclusions

Electro-mobility currently is the focus of consid-erable research effort, in both the publicly fundedand in the industrial domains. In this paper, weaddressed several key questions of electric flight,at the fundamental as well as at the system in-tegration level. We argued that three principlesshould be at the heart of any feasibility analysispertaining to airborne electrical systems: i) Ex-ergy (rather than energy) content is the parame-ter of choice to describe an energy carrier’s useful-ness for aviation. ii) This exergy density of the en-ergy carrier must be assessed in combination withthe power density acheived when using a partic-ular power conversion system, yielding the com-bined system’s Ragone metrics. iii) Several en-ergy carriers and power conversion systems whichseparately do not allow to sustain a given aircraftmission may be integrated into a suitable hybridelectrical power architecture capable of perform-ing this mission.

We further showed that the availabe state-of-the-art components for airborne electric systemsand in particular their future technology perspec-tives provide a promising starting ground to setup highly efficient hybrid electrical power archi-tectures (which we refer to as HEPAs), which –unlike conventional combustion engines – are notlimited by, e.g., the Carnot efficiency. We illus-

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0%

10%

20%

30%

40%

50%

60%

70%

0% 5% 10% 15% 20%

Fra

cti

on

al C

han

ge i

n R

an

ge (

-)

Fractional Change in Vehicular Aerodynamic Efficiency (-)

* ZFW is changed to TOGW if H = 0 and an all-battery energy source is utilized

Fig. 9 Nomograph showing impact of structural weight and aerodynamic efficiencies on range.

trated four distinct HEPA realizations, includingthe Universally-Electric System Architecture (U-ESA), which has been implemented in our recentpre-concept design CeLiner [5].

Based on the CeLiner analysis, we concludedthat the efficiency gains from electric flight can betruly optimized only if the entire solution spacespanned by both the physical system propertiesand aircraft configurations is explored. In general,the increased weight of an airborne U-ESA plat-form entails a sizeable reduction in mission range;we therefore studied in detail a hybrid system ofconventional turbo-engine and battery, which, de-pending on its degree of hybridization, allows forsignificant range extension, at the price, however,of losing the U-ESA’s zero inflight emission prop-erty. Interestingly, our analysis shows that there isa threshold DoH, i.e. a minimum fraction of theaircraft power demand that must be contributed bythe battery system, below which the hybridizationbenefits do not warrant the increased system com-plexity of the parallel gas-turbine and electric hy-brid. In order to mitigate the drawbacks of this

complexity, improved architectural solutions forhybrid energy and propulsion systems should re-ceive increased attention.

In light of this work, we conclude that progressin electric air transport is advancing at a fast pace,offering bright perspectives in the medium- tolong-term future. However, one must ensure theright tools are used to steer and quantify scientificdevelopments, and a reliable physics-based mod-elling approach should be applied to all compo-nents and sub-systems as well as the resulting in-tegrated aircraft concept.

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8 Copyright Statement

The authors confirm that they, and/or their company ororganization, hold copyright on all of the original ma-terial included in this paper. The authors also confirmthat they have obtained permission, from the copyrightholder of any third party material included in this pa-per, to publish it as part of their paper. The authorsconfirm that they give permission, or have obtained per-mission from the copyright holder of this paper, for thepublication and distribution of this paper as part of theICAS2012 proceedings or as individual off-prints fromthe proceedings.

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