Abstract Introduction

Downloaded from SAE International by Michele Trancossi, Tuesday, October 23, 2018

2018-01-1933

Published 30 Oct 2018

Energetic, Environmental and Range Estimation of Hybrid and All-Electric Transformation of an Existing Light Utility Commuter Aircraft Michele Trancossi Sheffield Hallam University Jose Pascoa Universidade Da Beira Interior Citation: Trancossi, M. and Pascoa, J., “Energetic, Environmental and Range Estimation of Hybrid and All-Electric Transformation of an Existing Light Utility Commuter Aircraft,” SAE Technical Paper 2018-01-1933, 2018, doi:10.4271/2018-01-1933.

Abstract

T

oday it is necessary to face the energetic, environmental, and safety-related issues of a significant industrial sector such as aeronautic one. It is a marginal contributor to today global GHG emissions (less than 3%), In any case, the associated impacts grows with the increase of air traffic with annual rate 5%. Consequently, aviation will need to face four fundamental problems for the future:

1. the overall impact of aviation is expected to grow up to 10÷15% of global GHG emissions by 2050; 2. the emissions of pollutants by commercial aviation affects the fragile atmospheric layers in the low stratosphere; 3. the increasing age of the flying fleet deals with increasing maintenance and safety issues; 4. the dependence on fossil fuels relates to problems of geopolitical instability and consequence volatility of prices.

Substantial innovations are expected for both reducing energy consumption and environmental impacts of aviation and reducing the age of the fleets. They mostly relate to the decrease of weights and the introduction of environmental friendly propulsion systems, such as hybrid and all-electric propulsion. This paper will produce an assessment of different propulsive systems according to the first law of thermodynamics and environmental impacts. It assumes a well-tested light transport/commuter aircraft as reference architecture and produces a comparative analysis of different green propulsion systems including all electric and hybrid against actual aircrafts. The analysis assumes that the electric or hybrid configurations may not increase the overall mass of the aircraft. Energy model has been reformulated for the different configurations and considers both an analytical model based on basic flight mechanics and a new formulation of the Breguet range equation, which has been specifically formulated for both hybrid and all-electric airplanes.

Introduction

Commercial aviation is a key sector of today society [1, 2]. It enables an unprecedented global connectivity and increases an effective global connectivity. The global growth of aeronautic industry is expected to continue over the next decades. Both Airbus and Boeing predict that passenger traffic alone will maintain a global average growth rate of around 4.7% per year up to 2036 (Figure 1). It must be remarked that ICAO estimates a higher growth rate of about 5% [3]. Consequently, the aircraft market is expected to grow even if at a lower rate (Figure 2). ICAO develops prediction models of aviation GHG emissions to 2040, and then extrapolates them to 2050 by assuming that 1 kg of jet fuel generates 3.16 kg of CO2. ICAO states that CO2 emissions in 2020 are assumed as the global aspiration of keeping the net CO2 emissions at this © 2018 SAE International. All Rights Reserved.

FIGURE 1 Global world annual traffic forecasts (Airbus [4])

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Aviation and Environmental Issues

Downloaded from SAE International by Michele Trancossi, Tuesday, October 23, 2018 2

ENERGETIC, ENVIRONMENTAL AND RANGE ESTIMATION OF HYBRID AND ALL-ELECTRIC TRANSFORMATION

3. reducing the weight of the aircraft by the adoption of lighter materials for the structures of the aircraft system; 4. increasing the aerodynamic efficiency of the aircraft by exploiting new aircraft designs such as the flying wing architecture with distributed propulsion and boundary layer suction; 5. reducing the dependency on fossil fuels, which will ensure the industrial sector from the geopolitical risks and consequent volatility of fuel prices? © SAE International

FIGURE 2 2017-2036 Market forecasts (Boeing [5])

FIGURE 3 ICAO global aviation emission scenario and uncertainty analysis [6]

More Electric, Hybrid and All-Electric Aircraft Technical Solutions


level (Figure 3). The second graph in Figure 3 shows the demand uncertainty effect on the fuel burn calculations [6]. The ICAO scenario and goals clearly state different directions for the future development of aviation: 1. greener, more efficient and less impacting propulsion systems; which include also hybrid and all-electric propulsion systems; 2. greener fuels which include innovative bio-fuels;

In addition, it must be remarked the necessity of c onsidering the fragile atmospheric layers in which the commercial aviation emissions are produced [7] and c onsequent radiative forcing. It is necessary to consider also the emissions near ground which can be evaluated in terms of both atmospheric pollution and significant acoustic emissions.

Many researchers are looking at hybrid and all electric propulsion as a mean for solving much of the environmental problems. In particular, these solutions seem to be the key solution for reducing the atmospheric and acoustic emissions. With respect to GHG emission they appear promising solutions through the ICAO goals of stabilizing the level of emissions at the values of 2020. Abe [8] analyses the problems and expectations that relate to hybrid vehicles. Emadi and Eshani [9] presents an effective review of electric systems for future aviation. Kim et al [10] presents the results of the NASA activity in the direction of a distributed propulsion hybrid flying wing aircraft. Bradley et al. [11] presents an energetic assessment of fuel-cells powered hybrid aircraft. Pornet at al. [12] present an effective assessment of the potential of fuel-battery hybrid narrow-body transport aircraft according to different design ranges for an entry-intoservice of 2035. Zhang et al. [13] analyses the different technical features of more electric and hybrid aircrafts. Metapon et al. [14] compare different energy management schemes for a fuelcell hybrid emergency power system of more-electric aircraft. Lieh et al. [15] presents an effective guideline for the design of hybrid propulsion system for unmanned aerial systems. Nahajagi [16] presents an exhaustive review of the industrial state of the art about technologies for more electric aircrafts. Williamson [17] discusses recent industrial developments in electric aircraft propulsion technology. In particular, he discusses Airbus Group's E-Fan, an electric propulsion element, E-Thrust, and hybrid electric aircraft research.

Aircraft Performance and Impacts Assessment Bejan and Siems [18] states the opportunity and necessity of assessing the second law exergy assessment of aircraft as a fundamental mean of greening and improving the performances of aircraft. Bejan [19] also states the opportunity of © 2018 SAE International. All Rights Reserved.



adopting tree shaped optimization methods in aircraft design. He demonstrates the opportunity of applying to aircraft optimization flow system architecture, coupling exergy analysis to determine the theoretical performance limit and thermodynamic optimization (or entropy generation). Rosen and Etele [20] have analyzed the exergy analysis method for aircrafts. They have discussed the differences between exergy analysis of both terrestrial and aerospace applications, and stated the effect of the ejection of exhaust gases with high exergy for an aerospace engine. They have also identified and discussed how it is possible to improve the thermodynamic efficiencies using exergy methods for aerospace applications. This activity has a fundamental role through a better comprehension of how exergy methods can be adopted to improve the performances and a better setup of the engines to different flight missions and operating conditions. In addition, Bejan et al. [21] demonstrates the key parameters for an effective aircraft efficiency looking at transport system as an element of natural evolution. Drela [22] has given a systemic formulation of aerodynamic phenomena in terms of thermodynamic magnitudes and has defined the general design drivers for aircraft energy and exergy optimization [23]. Arntz et al. [24] have produced a fundamental advancement of the Drela’s exergetic analysis of the aerodynamic behavior of an aircraft. Traub [25] assess the range and endurance of a battery powered aircraft. He clearly demonstrates the effects of increasing the battery capacity. He also demonstrates that a linear increase in capacity produces a nonlinear increase in required current. Seresinhe et al. [26] produce a comparative environmental assessment of traditional and more electric aircrafts. Baharozu et al. [27] compare traditional aircraft, more electric aircraft, and liquid hydrogen-fueled aircraft by using a multi criteria scoring method. This comparison method allows them to define a new aircraft concept that combines more electric and liquid hydrogen fueled concepts. The result is a new design that improves efficiency, cost, and environmental impact. In conclusion, the proposed aircraft concept for long range is an unavoidable choice for the future of the aviation sector to increase energy efficiency and decrease harmful environmental effects. Slivinsky et al. [28] present an effective design of a hybrid aircraft. They assess that the insufficient energy densities and specific energies of electrical storage devices are a potential showstopper to the development of all electric aircrafts as they induce severe weight and volume penalties. They focus on the increasing possibility of energy harvesting techniques, hybrid and more electric aircraft technologies. They consequently develop accurate parametric studies for traditional, electric and hybrid configurations. The final results define an effective model for determining range and endurance performances and their significant dependency on design parameters.

Forces Applied to an Aircraft Take off performance depends on the acceleration of the aircraft along the runway based on force equilibrium (Figure 4). The forces applied to the aircraft are: Thrust: T;

Weight : W = mg ; (1)

Drag force : D = 0.5 × C D × r × v 2 × S; (2)

Lift force : L = 0.5 × C L × r × v 2 × S; (3)

Friction force with the runway : F = m × (W - L ) (4)

The thrust of gas turbine or turbofan engines will be relatively constant during take-off and is linearly changing with velocity v Tv = Tstatic - K T × v (5)

A good assumption is to use the manufacturer's values for maximum static thrust for take-off calculations. The thrust of a propeller driven aircraft is found from the shaft power data for the engine and the propeller efficiency: T = ( Pshaft ×h P ) /v (6)

Propeller efficiency can be determined by mean of advance ratio J, that is the ratio of forward to rotational speed of the propeller, J = v / (nD), thrust coefficient, CT = T/(ρ n2 D4), and torque coefficient, CQ = Q/(ρ n2 D5). The propulsive efficiency results:

h P = ( J /2p )( CT /CQ ) (7)

At v = vR the efficiency will be in the range 50% to 80% depending on the type of propeller and the thrust value at this point will be easy to obtain. In practice, the thrust obtained throughout the take-off roll is roughly constant. The thrust of a gas turbine aircraft can be obtained by manufacturers’ charts. They usually quote static sea level thrust (Tstatic). If this is the case, then estimates of thrust at altitude or high speed must be made. Variation of static thrust with altitude can be approximated as n

T (h ) æ r (h ) ö =ç ÷ , (8) T ( 0 ) çè r ( 0 ) ÷ø

where n varies between 0.7 at sea level and 1.0 at cruise conditions in the stratosphere. Thrust variation with speed is approximately linear between static and cruise conditions and is expressed by equation (2). FIGURE 4 Equilibrium of an aircraft moving on the ground

The definition of a satisfactory first and second law models requires the definition of the forces, which are applied on the aircraft, and the Newtonian equation of motion. © 2018 SAE International. All Rights Reserved.


Methodology Aircraft Operations

3



The drag coefficient CD is variable in terms of CL C D = C D 0 + K L × C , where K L = 1/ (p × AR × e ) . (9)

The friction coefficient μ for a standard tarmac runway is 0.02. Take Off Model During take-off run the aircraft accelerates along the runway according to equation (4).

dv /dt = (T - D - F ) /m (10)

where dv/dt is the acceleration along the runway and m is the mass of the vehicle. It must be remarked that wings during take off are in maximum lift configuration and CD = CDmax ≃ const. The vehicle will accelerate until it reaches a safe flying speed vR, which is

v R = 1.2v stall = 1.1v mincontrol . (11)

The safe flying speed (or rotation speed), vR, is a critical factor in both safety and determining take-off performance and is determined at the maximum achievable lift coefficient CLmax:

v stall = W / ( 0.5 × C L max × r × S ) (12)

The acceleration during takeoff operation is given by equation (1) and the velocity can be predicted in any point of the runaway by integrating equation (1) over time. The acceleration is almost constant during take off operations and this average acceleration corresponds to the acceleration that is assumed when v = v R / 2 . Assuming a constant acceleration, equation (13, 14) can be determined:

v R = a × t R (13)

s1 = 0.5 × a × t R2 = 0.5 × v R2 /a (14)

From the rotation point, the end of the runway can be defined by the requirement to clear a 35ft (10,7 m) obstacle at the end. During rotation eventual residual thrust excesses are absorbed by the lift induced drag during the vey preliminary climb. Acceleration reduces up to zero and the distance between rotation and clearance points is

The climb angle dh/dt is expressed by dt / dh dh (T - D ) × V Þ = . (18) V dt W An accurate management of the necessary excess power allow optimizing the climb rate. The Specific Excess Power Ps:

g = sin g =

Ps =

(T - D )V (19)

M×g It is specifically required to climb or to overcome energy losses during motion. Level Flight When the aircraft is in equilibrium at steady level flight, it is

T = D (20) For level flight analysis, specific Excess Power generates a climb rate, but it is necessary during level flight for reinstating the equilibrium of the aircraft in case of disturbances and manoeuvring.

Descent It is obvious that during descent T < D (Figure 7). Consequently, it can be expressed by

ìT - D = -W sin g (21) í î L = W cos g

The optimization of descent allows minimizing fuel usage and maximising the distance travelled. To minimise fuel consumption, the thrust should be reduced to a minimum (T=0). To maximise distance travelled then the aim will be to have the smallest descent angle with no thrust, that is, to find

FIGURE 5 Forces on the aircrafts during climb


2 L

35 10.67 ( ft ) = ( m ) (15) tan g rot tan g rot The initial climb gradient can be calculated by equation (13).

s2 =

Climb Figure 6 shows the dynamic model of the aircraft during climb. The vehicle is assumed to climbing at a constantangle (γ), constant forward velocity (v), and constant rate (dh/dt). During climb, thrust T exceeds drag D. The balance of forces is:

FIGURE 6 Forces on the aircraft during descent

ìT - D = W sin g (16) í î L = W cos g

g sin g =


The relation between forces and the angle of climb γ can be determined by assuming γ small T -D . (17) W © 2018 SAE International. All Rights Reserved.



the best glide angle. The descent angle γ can be considered small and consequently it can be obtained the gliding condition:

L = W Þ sin g g =

D 1 . (21) = W L/D

The landing run can be calculated using a similar method to the take off distance. The aim is the minimization of the landing distance, also for safety reasons. Assuming that the touch down speed vTD is almost equal to the stall speed of the aircraft in landing configuration, it is achieved by a pitch maneuver during the flare portion of the approach which increase drag an decelerate the aircraft to minimum flying speed. The deceleration on the landing roll from vTD to v0 will be accomplished by braking and reverse thrust. This can be solved by the average acceleration approach that was used to estimate the take-off roll.

dv -T - D - F = (22) dt m

5

Minimum power condition can be determined by differentiating equation (25) with respect to speed v:

3 K L ×W 2 × CD0 × r × S × v 2 = (27) 2 0.5 r × S × v 2

By substituting equation (9) into (27), it is possible to determine CD for minimum power (during cruise flight):

C Di = 3C D 0 = KC L2 (28)

Energy Balance of a Traditional Aircraft The basic energy balance of a traditional aircraft during its operations is presented in Figure 8.

Ein = E prop + Elosses (29) The energy given by the fuel is

The negative acceleration or deceleration value will be based on maximum braking friction coefficient and the reverse thrust (when available).

Power Balance

The balance of energy components Ein=E out allows obtaining the general equation that applies during all the flight operations, which is expressed by equation (31):

The first law analysis is performed in terms of power. The power required to keep an aircraft in steady level flight is

P = T × v = D × v = 0.5 × C D × r × v 3 × S (23)

By substituting into equation (23) CD from equation (9) it results

(

)

P = 0.5 × C D 0 + K LC L2 × r × v 3 × S

dEin = dM fuel × LHV fuel (30) dE prop = dEin ×htot = dM fuel × LHV fuel ×htot = T × v × dt

æ T ö vx ç ÷ dW f - Dv x dt - Fv x dt - Wv y dt = DE (31) dW dt / f è ø

Assuming that the fuel consumption Wf reduces the mass of the aircraft –W and that Trust Specific Fuel Consumption is equal to

TSFC =

dW f /dt , (32) T

For level flight, it is L = W, and consequently it results

D C D C D 0 + K LC L = = (25) L CL CL

The minimum D/L condition can be found by differentiating D/L:

v min drag =

FIGURE 8 Schema of an all electric aircraft

CD0 (26) KL

CL =

By substituting (26) into the steady flight equation, the airspeed for minimum drag assumes the value


The required lift at the various airspeeds should be constant. The minimum drag speed can be determined by

FIGURE 7 Energy balance of a traditional aircraft

W KL 0.5 r S C D 0

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æ W 2 ö -1 P = ( 0.5 × C D 0 × r × S ) × v + ç K L ÷ × v (24) è 0.5 r S ø 3



Equation (31) becomes

Vx -dW D F DE - dx - dx - dy = = dhe (33) TSFC W W W W During horizontal cruise it becomes Vx -dW D - dx = dhe 0 (34) TSFC W W and consequently Vx L dW (35) TSFC D W The range is obtained by integrating equation (35): dx = -

æ -Vx L dW ö ç ÷ (36) Winitial è TSFC D W ø It results equation (37) R=

ò

W final

lift to drag ratio. Consequently, the range can be obtained by considering the energy which is stored in the battery system:

D E *mbatthtot L R®R= (39) L W D The range R results consequently E *mbatthtot = W

R = E*

mbatt L htot (40) mg D

Consideration about Batteries Charge and Discharge The actual model of all electric aircraft requires an effective analysis of battery charge and discharge modes. Batteries discharge is simply described by Peukert’s equation (41). t = C /i nd (41)

Vx L æ Winitial ö Vx L æ minitial ö ln ç ln ç ÷= ÷ (37) TSFC D è W final ø TSFC D è M final ø It is the classic Breuget range equation, which estimates the range of an aircraft in still air. However, the assumption of constant L/D and constant v/TSFC may not be too accurate in practice. Keeping constant the ratio L/D with the changing weight the aircraft would need to drift up in altitude so that a constant angle of attack can be maintained. It may also be required that the aircraft change speed to maintain a constant V/TSFC. Consequently, a detailed estimation of long-range cruise conditions may require a numerical summation over short segments where the assumptions are accurate. In presence of headwind (V = Vx - V0) or tailwind (V = Vx + V0) the energy balance becomes

where t is the time in hours, i is the discharge current (amperes), and C is the battery capacity in ampere hours and n is a discharge parameter dependent on the battery type and temperature. It must be observed that nd changes during the lifecycle of a given battery as it ages and cycles such that capacity usually diminishes. In order to account the effect of the discharge rate, it becomes:

æ T ö v xç ÷ dW f - D (Vx ± V0 )dt - F (Vx ± V0 ) dt -WVy dt = DE è dW f /dt ø

Energy Balance of Hybrid Aircraft

R=

and the range formula can be modified as: R =

(Vx ± Vv0 ) TSFC

L æ Winitial ö (Vx ± V0 ) L æ minitial ö ln ç ln ç ÷= ÷ (37) D è W final ø TSFC D è m final ø

Energy Balance of All Electric Aircraft An electric aircraft makes use of batteries for energy storage and an electric motor for propulsion (Figure 8). The amount of electrical energy an aircraft can use effectively is a function of battery specific energy, battery mass and propulsion efficiency. An all-electric aircraft has a constant mass during flight and that it do not benefit of the reduction of weight, which is consequent to the fuel consumption.

E = E *mbatthtot (38)

where E*mbatt is the electrical energy that can be stored. The amount of thrust which is needed is proportional to the

nd

tr æ C ö ç ÷ , (42) i n è Rt ø in which tr is the discharge time in hours. For a battery, the output power may be estimated as (where V is volts).

t=

PB = Ri = V

1/ nd

C æ tr ö ç ÷ Rt è t ø

(43)

Discussion on hybrid aircrafts will require analyzing three different configurations. In the case of series hybrid and FC hybrid configuration it can be possible to account the energy balance of the system in a similar way to the one which is used for traditional aircrafts powered by fossil fuel. It is necessary to compute the system efficiency by considering the more complex layouts. Series and FC Hybrid Aircraft In this case, it is necessary to compute adequately the energy balance according to Figure 9 and 10. D R, (44) L where Wav is the average weight of the aircraft. Consequently, the aircraft endurance can be computed on the basis of the initial level of charge of the batteries c%:

W fuel × LHV fuel ×htot + E * × mbatt ×htot ,el = Wav

R=

æm ö vx L m L × × ln ç initial ÷ + c % × E * × batt × ×htot . (45) TSFC D m mg D è final ø © 2018 SAE International. All Rights Reserved.



FIGURE 11 Parallel hybrid configuration



FIGURE 9 Series hybrid configuration

7

FIGURE 10 Fuel Cell Hybrid configuration

Analysis of the Emissions © SAE International

The analysis of the emissions is a direct function of the mass of fuel. For each standard combustible, which is used in propulsion, it can be possible to determine emission factors that allow determining the missions in terms of both GHG gasses and other pollutants. In addition, it is necessary to consider the radiative effect [28] that allows accounting the height at which emissions occurs.

Hybrid Parallel Configuration In the case of a series hybrid configuration much more complex considerations must be necessarily performed. Different optimization methods and algorithms have been developed. In particular for a solution of the problem it is necessary to optimize the times and the electric and ICE powers. It must be necessary to define the energy absorption by the electric generation system. Those factors present a fundamental effect on the results. In order to assess an effective model based on the former theory, a coefficient of usage ε is introduced. The energy produced before the generator and the power transmission is given by

Eeng = E fuel ×heng (46) The two components of the power system are

Ein ,mot = e Eengh genhbathinvhmot (47)

Ein ,eng = (1 - e ) Eeng (48) In this case we can assume that the propulsive energy is

E prop = ( Ein ,mot + Ein ,eng ) ×htransm ×h prop (49)

In the case in which different transmission and propulsion systems it becomes:

E prop = Ein ,mot ×htrans ,mot ×h prop,mot + Ein ,eng ×htrans ,eng ×h prop,eng

In this case it can be possible to use equation (45) to determine the range of the aircraft. © 2018 SAE International. All Rights Reserved.

Solution Comparison between the Energy Efficiencies of the Different Systems It is possible to realize a preliminary comparison between different systems in terms of relative efficiency. In particular, the much larger number of considered configurations allows extending the angle of evaluation performed by Epperle. They have been evaluated according to the former discussion. Reference efficiencies, which have been assumed for the calculation of the different propulsive models, have been reported in Table 1. The efficiencies of different systems have been computed into Figure 12. In particular, the hybrid systems account a ratio ε = 0.5 between hybrid and combustion unit.

Reference Aircraft It has been assumed an interesting twin-engine aircraft such as Britten Norman BN2 Islander [29]. Main technical data from the producer website have been reported in Appendix. It is a 10 seater light utility aircraft and regional transport aircraft, which has been designed to have with low costs and easy maintenance. Designed during the ‘60s is still in production for its own specific characteristics of operational flexibility and short take off and landing. The high wing lift around 85.4 kg/m2 ensures excellent STOL (short Takeoff and Landing) capability with a take off run of 140 m, a pretty low cruise speed of 257 km/h, a very low stall speed (64 km/h), an excellent rate of climb 295 m/min [29].



TABLE 1 Reference efficiencies for the evaluation of different

FIGURE 13 Britten Norman BN2 Islander drawings

components

Cycle efficiency Internal combustion engine

36%

Turboprop

50%

Turbofan

50%

Battery

100%

Fuel Cells

60%

Hybrid series ICE

42%

Hybrid series Turboprop

55%

Hybrid parallel ICE

40%

Hybrid parallel Turboprop

52%

Generator

95%

Controller

98%

electric motor

95%

Gearbox

98%

Propeller

80%

Turbofan Fan and Nozzles

60%


Reference component efficiency


FIGURE 12 Efficiency of different propulsion systems

Performances, dimensions and weights of the aircraft allow identifying very interesting perspectives for electrification [30]. MTOW in takeoff configuration is 2994 kg and a maximum payload of 774 kg without tip tanks. It is equipped by two Lycoming IO540 [31] with a max power of 195 kW and cruise at 65% power (126.75 kW). The twin-engine propulsion system has an overall dry mass of 400 kg. The maximum attainable payload is 776 kg will also allow a good battery load. The empty weight is around 1670 kg (16380 N) against a MTOW of 2994 kg (29370 N). It is consequently possible to have an available mass of about 1720 kg for electric storage, propulsion and transported payload.

Components of Energy Systems Battery Different battery options could be available [32]. An accurate analysis of commercially available battery packs

Energy Energy Density Density Energy Weight Volume by weight by size stored Battery type

kg

cm3

Wh/kg

Wh/cm3 Wh

SLA (3x12V 10Ah) 10,00

3270

36

0.11

360

NiMH (36V10Ah)

2430

65

0.15

357,5

5.50

Poly Li-Ion (8Ah) 1.75

1340

170

0.23

297,5

Poly Li-Ion (10Ah) 2.15

1613

170

0.23

365,5

has been performed. In particular, a preliminary comparison between different cells and solution is presented in Table 2. A leading battery seller [33] allows producing an effective comparison of the data. In particular it is evident that Li-Ion batteries are the best solutions. Considering the actual market leader solution Tesla 2017 Model 3 [40] batteries present even better performances: •• 444 cells (74 parallel and 6 series) •• 5.3 kWh, 24 V •• 56.45 lbs (25.6 kg) -> 207 Wh/kg Electric Motors Looking at top global electric motors manufacturers, the best industrial solution appears to be YASA 750 R [34] motor (Table 3). © 2018 SAE International. All Rights Reserved.



TABLE 2 Comparison between different battery types [33]



TABLE 3 YASA 750 R motor technical characteristics

TABLE 4 YASA 750R technical data

(from [34])

Parameter

Electrical

Mechanical

Peak Current (10 secs)

600 ARMS

Continuous Current

300 ARMS

Total volume 7 L

Input Voltage Range 50 - 400 VDC

50 ÷ 400 VDC

100 kW

Axial Length

98 mm

Efficiency

96%

Peak power 700V

200 kW

Diameter

368 mm

Dimensions

84 x 251 x 277 mm

Continuous power

Up to 70 kW*

Coolant

Oil

Mass

5.0 kg

Maximum Speed

3250 rpm

Flow rate

20 L/min

Peak efficiency

> 95%

Fluid volume 0.6 L

Peak torque

790 Nm

Total weight

Continuous torque

400 Nm

Peak Power 350V

37 kg



9

Total Volume

5.8 litres

Volume excluding HV terminal area

5.1 litres

* Subject to drive cycle and thermal conditions


FIGURE 14 YASA 750 R drawings (from [34])

FIGURE 15 Velocity and voltage against power


A YASA 400V 600A Controller is an adequate solution for applications such as hybrid and electric vehicles. They allow achieving low space and weight with a peak power density up to 40 kVA / L. The 400V, 600A type controller is specifically designed for YASA P400 and 750 type motors and generators with peak power up to 200kW. The possibility of an effective use of the YASA 400V 600A controller can be verified by mean of the graph of power against speed for the specific aircraft. It is reported in Figure 15. Considering the vehicle characteristics, it is evident that the indicated motor can fulfill the need for the specific aircraft even if it can need some aerodynamic improvements in order to reduce the drag. Otherwise, the maximum safe continuous power that is 70 kW for the specific motor will require reducing the cruise speed to 210 km/h (it can be a suitable option for the all-electric version). It must be also evaluated the option of producing an additional thrust which can be evaluated in the case of a hybrid version. For environmental assessment of electricity emissions the EU including UK electricity emission factor is 0.46 kg(CO2e)/kWh [42, 43]. It is fundamental to refer to NASA X57 project [35, 36]. It adopts P2006T fuselage for all-electric propulsion with a new wing concept with diffused propulsion. The first X57 run prototype allowed evidencing one of the key problems of aeronautics electrification that deals with battery thermal management. During 2016 tests, a battery cell was shorted and the overheating spread to other cells. This unwanted effect requires redesigning the battery modules. The design of the battery system must account those problems.

Hybrid Aircraft Configurations Engine It is necessary to define the options for the hybrid versioning. In particular, it is necessary to adopt a generator, which has not less than 140 kW of effective power. An interesting choice could be a Rolls-Royce M250B-17 turboprop [36, 37] with a max power of 420 kW and a mass of 93kg. An internal combustion engine alternative could be a well tested piston engine such as the Continental turbocharged Diesel cycle CD-300 [39] with a max power of 221 kW and a mass of 248 kg. It is evident that the turboprop solution presents a much lower power to weight ratio. For a parallel hybrid configuration a Rotax 912 [41] turbocharged engine has been considered for its favorable power/mass ratio (100 kW/84kg). The definition of the different configurations can be now performed. According to the above discussion, the empty mass without motors of the aircraft is 1270 kg. Max take off mass is 2994 kg, and 1724 kg can be used for the transformation. Electric Generation The electric generation system is assumed to be constituted by referring to the PCU Assembly




FIGURE 19 Toyota Hybrid system schematics [44]



FIGURE 16 Continental CD300 engine [39]

FIGURE 17 Allison/Rolls Royce 250 turboprop drawing (from manual) TABLE 5 Toyota 4th generation PCU (Power Control Unit

Max Power in

132 kW

Max Electric Power Out

90 kW

AC-Inverter

Integrated

Generator AC-Inverter

Integrated

Boost Converter

Integrated

DC/DC Converter

Integrated

Cooling

Liquid-cooled

Volume

6.8 Liters

Weight

11.9 kg

25 mm2 [45] it can be assumed a mass of 14.4 kg. 26 kg are prudentially estimated for necessary accessories. The losses for cabling are considered included into the energy conversion factor.


FIGURE 18 Rotax 91 IS [41]

schema by TNGA (Toyota New Global Architecture), which is reported in Figure 19 [44]. Two Toyota power units can be estimated for the above propulsion architecture, to allow an adequate multiplicity and safety with a total mass of the PCU of about 24 kg. Cabling and Accessories Assuming to use 6 mm2 cables for power distribution it can be estimated a length of around 50 m. Adopting a standard PVC insulated cable in accordance with ISO 6722 with a resistance of 0.743 Ω/km and an area of

Final Considerations Redundancy has not been considered at this stage, because structural weight reductions are not considered at this stage. In particular assuming to equip the aircraft with two YASA 750 electric motors and relative controls and cabling, the available mass for electric generation, storage and payload will be around 1640 kg. After a large research on available batteries, it has been adopted Tesla Model 3 batteries, because of their outstanding performance and their actual market disposability. New battery models are expected on the market but their commercialization has not started. Four different configurations have been considered against the original one. Two different calculation have been performed, one by the traditional equation of flight by considering the reference mission used by Stoll and Veble Milic [46] and by the modified range equation which has been obtained in this research activity. The results has lead in a difference of not more that 5%, which is acceptable for the proposed research. © 2018 SAE International. All Rights Reserved.



Data) [44]



11

TABLE 6 Comparison of different configurations of the reference aircraft.

BN 2 Islander

All- electric

Hybrid Series (CD300)

Hybrid Series (RR250)

Hybrid parallel (Rotax 915)

Engine

400

0

176

120

168

kg

Electric Generator

0

100

100

50

kg

PCU Cables and accessories

0

50

50

50

50

kg

Battery packs

0

50

27

27

24

n.

Energy Stored

0

265

143,1

143,1

127,2

kWh

Battery packs mass

0

1280

691,2

691,2

614,4

kg

Motors and accessories

30

145

145

145

145

kg

Fuel

549

0

230

300

320

kg

Empty weight

1667

1270

1270

1270

1270

kg

Payload

774

254

436,8

422,8

422,6

kg

Cruise speed

257

208

208

208

257

km/h

Electric range

0

350

200

200

120

km

0

516

516

892

km


Hybrid range Fuel range

1400

0

0

0

0

km

Max range

1400

350

716

716

1012

km

Endurance

5h 25'

1h 42'

3h

3h

4h 35'

h

CO2e emissions/km

1,239

0,348

1,107

1,416

1,057

kg

CO2e emissions/km variation

100,0%

-71,9%

-37,9%

16,0%

-12,9%

%

Results The results have been evaluated for the different configurations, which have been considered. Considering both analytical and range equation, the most conservative results have been reported. They have been reported in Table 4. The results show clearly that hybrid and all electric porting of existing aircraft models can be performed. The adoption of today technologies still requires an appreciable reduction of the maximum available range. It can be also observed that all the presented solutions reduce the CO2 emissions with major environmental benefits. Future available technologies are expected to produce an effective increase in terms of battery energy density and discharge speed. The current technology is suitable for small Light aircraft, but not for commercial aviation. In order to power even small commercial aircrafts such as Islander a dramatic improvement in battery technology would be required. Today battery technology presents energy densities in a range between 150 and 210 Wh/kg. For an effective use on all electric aircrafts the energy density must be increased by a factor not lower than 3.

Conclusions Practical all-electric airplanes are today limited to small vehicles up to 2 passengers. They still have problems related to the low energy density of available batteries, and consequent limitations in terms of range and endurance. This paper analyses if electrification is suitable for larger 3 tons class light regional transport aircrafts. It reaches this goal by defining an energy model that considers the most significant flight phases. It does not consider actually fuel cells © 2018 SAE International. All Rights Reserved.

hybrid aircrafts because of the large literature on the specific technical solution. It applies the model to a well-known regional transport aircraft and analyses different electric configurations: all electric; hybrid series powered by a diesel engine, hybrid series powered by a turbine; hybrid parallel with two small engines. The results show clearly that today market ready batteries are still not ready for commercial aviation. Well performing electric motors that can satisfy commercial aviation requirements are available. On the other side, batteries are still not adequate for this role, at least in all electric configurations. They still perform adequately just for hybrid propulsion. It must be remarked that the low energy density of available batteries requires a reduction of the payload to achieve an adequate range. This problem is much more evident in the all-electric configuration, but also hybrid configurations are affected by the same problem. More realistically, this factor would have to be in the order of 6 to 10 to attract commercial interest for larger (regional) aircraft. In this context, it must be observed that considered energy reserves for all electric aircraft are limited to 10% and battery-aging phenomena are not considered. This assumption means that only few minutes of flight are ensured for holding at the destination airport, delays, and deviations. The results show that 3 over 4 configurations reduce greenhouse gasses emissions. This reduction is much more significant for all-electric configuration and is much lower in the case of hybrid aircrafts. In particular, greenhouse emissions can result higher for a series hybrid configuration with a turbine electric generator. This paper presents an effective guideline for energetic, environmental and range assessment. On the other side, it opens a series of questions, which have to be addressed in the direction of an effective implementation of all electric aircrafts:



•• a complete lifecycle evaluation of energy balances and environmental footprint of electric aircraft, which includes also manufacturing and end-oflife dismantling; •• an effective estimation of the effects of the needs of raw materials for the production of the huge battery amount which is required by the electrification of commercial aviation; •• estimation of infrastructures and investments, which are necessary for battery recharging and eventual replacement; •• definition of both the required levels of safety and the necessary certification rules.

References 1. Airbus, “Global Market Forecast 2017-2036 - Growing Horizons,” Airbus Industries, 2017, ISBN:978-2-9554382-2-6, https://www.airbus.com/content/dam/corporate-topics/ publications/backgrounders/Airbus_Global_Market_ Forecast_2017-2036_Growing_Horizons_full_book.pdf, accessed June 2018. 2. Boeing, “Current Market Outlook 2017-2036,” Boeing Corp., 2017, accessed June 2018, http://www.boeing.com/resources/ boeingdotcom/commercial/market/current-marketoutlook-2017/assets/downloads/2017-cmo-6-19.pdf. 3. ICAO, “Agenda Item 17: Environmental Protection Present and Future Trends in Aircraft Noise and Emissions,” International Civil Aviation Organization, Montreal, 2013. 4. Rädel, G. and Shine, K.P., “Radiative Forcing by Persistent Contrails and Its Dependence on Cruise Altitudes,” Journal of Geophysical Research: Atmospheres 113(D7), 2008. 5. Brasseur, G.P. and Gupta, M., “Impact of Aviation on Climate: Research Priorities,” Bulletin of the American Meteorological Society 91(4):461-464, 2010. 6. VV.AA., “Aviation’s Contribution to Climate Change,” ICAO, Environmental Report 2010, Vol. 1. 7. Emadi, A. and Ehsani, M., “Electrical System Architectures for Future Aircraft,” SAE Technical Paper 1999-01-2645, 1999, doi:10.4271/1999-01-2645. 8. Kim, H.D., Brown, G.V., and Felder, J.L., “Distributed Turboelectric Propulsion for Hybrid Wing Body Aircraft,” 2008 International Power Lift Conference, Royal Aeronautical Society, London, July 22-24, 2008. 9. Bradley, T., Moffitt, B., Parekh, D., Fuller, T. et al., “Energy Management for Fuel Cell Powered Hybrid-Electric Aircraft,” 7th International Energy Conversion Engineering Conference, 2009, 4590. 10. Pornet, C., Kaiser, S., Isikveren, A.T., and Hornung, M., “Integrated Fuel-Battery Hybrid for a NarrowBody Sized Transport Aircraft,” Aircraft Engineering and Aerospace Technology, An International Journal 86(6):568-574, 2014.

11. Zhang, H., Saudemont, C., Robyns, B., and Petit, M., “Comparison of Technical Features between a more Electric Aircraft and a Hybrid Electric Vehicle,” Vehicle Power and Propulsion Conference, 2008. VPPC’08, IEEE, Sept. 2008, 1-6. 12. Motapon, S.N., Dessaint, L.A., and Al-Haddad, K., “Comparative Study of Energy Management Schemes for a Fuel-Cell Hybrid Emergency Power System of MoreElectric Aircraft,” IEEE Transactions on Industrial Electronics 61(3):1320-1334, 2014. 13. Lieh, J., Spahr, E., Behbahani, A., and Hoying, J., “Design of Hybrid Propulsion Systems for Unmanned Aerial Vehicles,” 47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, Aug. 2011, 6146. 14. Naayagi, R. T., “A Review of More Electric Aircraft Technology,” IEEE International Conference on Energy Efficient Technologies for Sustainability (ICEETS), 2013, Apr. 2013, 750-753. 15. Williamson, M., “Air Power the Rise of Electric Aircraft,” Engineering & Technology 9(10):77-79, 2014. 16. Bejan, A. and Siems, D.L., “The Need for Exergy Analysis and Thermodynamic Optimization in Aircraft Development,” Exergy, An International Journal 1(1):1424, 2001. 17. Bejan, A., “Constructal Theory: Tree-Shaped Flows and Energy Systems for Aircraft,” Journal of Aircraft 40(1):4348, 2003. 18. Rosen, M.A. and Etele, J., “Aerospace Systems and Exergy Analysis: Applications and Methodology Development Needs,” International Journal of Exergy 1(4):411-425, 2004. 19. Bejan, A., Charles, J.D., and Lorente, S., “The Evolution of Airplanes,” Journal of Applied Physics 116(4):044901, 2014. 2 0. Drela, M., “Power Balance in Aerodynamic Flows,” AIAA Journal 47(7):1761-1771, 2009. 21. Drela, M., “Design Drivers of Energy-Efficient Transport Aircraft,” SAE Int. J. Aerosp. 4(2):602-618, 2011, doi:10.4271/2011-01-2495. 22. Arntz, A., Atinault, O., and Merlen, A., “Exergy-Based Formulation for Aircraft Aeropropulsive Performance Assessment: Theoretical Development,” AIAA Journal 53(6):1627-1639, 2014. 2 3. Traub, L.W., “Range and Endurance Estimates for Battery-Powered Aircraft,” Journal of Aircraft 48(2):703707, 2011. 2 4. Seresinhe, R., Lawson, C., and Sabatini, R., “Environmental Impact Assessment, on the Operation of Conventional and More Electric Large Commercial Aircraft,” SAE Int. J. Aerosp. 6(1):56-64, 2013, doi:10.4271/2013-01-2086. 25. Baharozu, E., Soykan, G., and Ozerdem, M.B., “Future Aircraft Concept in Terms of Energy Efficiency and Environmental Factors,” Energy 140:1368-1377, 2017. 2 6. Sliwinski, J., Gardi, A., Marino, M., and Sabatini, R., “Hybrid-Electric Propulsion Integration in Unmanned Aircraft,” Energy 140:1407-1416, 2017.




27. IPCC, “Climate Change 2013: The Physical Science Basis. Contribution of Working Group I to the Fifth Assessment Report of the Intergovernmental Panel on Climate Change,” Cambridge University Press, Cambridge, and New York, 2013, 1395-1445, published Jan. 31, 2014. 2 8. VV.AA, Islander Owners Handbook (UK, Bitten Norman, 1975). 29. EASA, “EASA.A.388 Britten Norman BN2,” European Aviation Safety Agency, 2011, https://www.easa.europa. eu/documents/type-certificates/aircraft-cs-25-cs-22-cs23-cs-vla-cs-lsa/easaa388. 3 0. VV.AA, “Britten Norman BN2 Islander Brocure,” Britten Norman, UK, 2011. 31. VV.AA., “Operator’s Manual Lycoming O-540, IO-540 Series,” Approved by FAA, Lycoming, 2006, https://www. lycoming.com/sites/default/ files/O%20%26%20IO-540%20Oper%20Manual%206029710.pdf. 32. Kim, J.G. et al., “A Review of Lithium and Non-Lithium Based Solid State Batteries,” Journal of Power Sources 282:299-322, 2015. 33. VV.AA., “Battery Space Catalog,” retrieved June 2018, http://www.batteryspace.com/. 3 4. YASA, “YASA 700 R E-Motor Product Sheet,” retrieved June 2018, http://www.yasa.com/wp-content/ uploads/2018/01/YASA-750-Product-Sheet.pdf. 35. Falck, R.D., Chin, J., Schnulo, S.L., Burt, J. M. et al., “Trajectory Optimization of Electric Aircraft Subject to Subsystem Thermal Constraints,” 18th AIAA/ISSMO Multidisciplinary Analysis and Optimization Conference, 2017, 4002. 36. Papathakis, K.V., Ehmann, D., Sessions, A., and Burkhardt, P., “A NASA Approach to Safety Considerations for Electric Propulsion Aircraft Testbeds,” 2017. 37. VV.AA, “M250 First Network Directory,” Rolls Royce, UK, 2017, accessed June 2018, https://www.rolls-royce. com/~/media/Files/R/Rolls-Royce/documents/customers/ defence-aerospace/M250%20FIRST%20network%20 directory%202015.pdf. 38. Pegors, D., “Advanced Allison Small Turboprop Engines,” SAE Technical Paper 871055, 1987, doi:10.4271/871055. 39. VV.AA, “FAA Lycoming IO-540 Series Type Certificate,” 2006, Retrieved June 2018, http://rgl.faa.gov/Regulatory_ and_Guidance_Library/rgMakeModel.nsf/0/ffae5a2bb550 6dcc8625747a00650001/$FILE/1E4.pdf. 4 0. Bower, G., “GM Versus Tesla: Bolt EV And Model 3 Battery Packs Compared,” InsideEVs, 2018, accessed June 2018, https://insideevs.com/gm-versus-tesla-bolt-ev-teslamodel-3-battery-packs-compared/. 41. VV.AA, “Rotax 915 915|141 hp (iS/iSc) Technical Description,” BRP-Rotax GmbH & Co KG, 2017, https:// www.flyrotax.com/files/Bilder/Produkte%20Rotax/ Datasheets/Data_sheet_915%20iS_iSc_A4_18.01.2018.pdf. 4 2. VV.AA., “ISCC 205 GREENHOUSE GAS EMISSIONS Version 3.0,” ISCC GMBH, DE, 2016, Valid from: Aug. 9, 2016 (Date of Commission Implementing Decision (EU) 2016/1361), https://www.iscc-system.org/wp-content/ uploads/2017/02/ISCC_205_GHG_Emissions_3.0.pdf. © 2018 SAE International. All Rights Reserved.

13

4 3. VV.AA., “EPA Emission Factors for Greenhouse Gas Inventories,” US Environmental Protection Agency, 2018, https://www.epa.gov/sites/production/files/2018-03/ documents/emission-factors_mar_2018_0.pdf. 4 4. Warner D.M. et al., “Advanced Strong Hybrid and Plug-In Hybrid Engineering Evaluation and Cost Analysis,” CARB Agreement 15CAR018, Principal Investigators: Al Steier, Munro & Associates, and Alan Munday, Ricardo Strategic Consulting, 2017. 45. VV.AA., “Flexible Cables and Cords,” Anixter Ltd., 2017. 46. Stoll, A.M. and Veble Mikic, G., “Design Studies of ThinHaul Commuter Aircraft with Distributed Electric Propulsion,” 16th AIAA Aviation Technology, Integration, and Operations Conference, 2016, 3765.

Definitions/Abbreviations α - Angle of attack [°] γ - Angle of climb/descent [°] ηp - Propulsive efficiency [−] ηtot - Propulsive efficiency [−] μ - Friction coefficient [−] ρ - Density [kg/m3] C - Battery capacity [Ah] CD - Drag coefficient [−] CL - Lift coefficient [−] CQ - Torque coefficient [−] CT - Trust coefficient [−] D - Drag force [N] E - Energy [J, kWh] Eeng - Energy output by the engine [J, kWh] Ein - Energy input [J, kWh] Emot - Energy output by the electric motor [J, kWh] Ein,eng - Energy input by engine [J, kWh] Ein,mot - Energy input by electric motor [J, kWh] Elosses - Energy dissipations [J, kWh] Eprop - Propulsion energy [J, kWh] E* - Energy density [J/kg; kWh/kg] J - Advance ratio [−] F - Friction force [N] K L - Drag corrective coefficient [−] L - Lift force [N] LHVfuel - Lower heating value of fuel [J/kg] P - Propulsive power [W] Pb - Battery output power [W] Pshaft - Shaft Power [W] R - Range [km] S - Surface [m2] T - Trust [N]



Tstatic - Static thrust [N] TSFC - Trust Specific Fuel Consumption [W/N] Tv - Dynamic thrust [N] W - Weight [N] a - Acceleration [m/s2] nd - Battery discharge coefficient [−] n - Shaft rotation speed [rpm; rad/s] g - Gravitational acceleration [9.81 m/s2] i - Discharge current [A]

m - Mass [kg] mbatt - Mass of batteries [kg] mfinal - Final mass of the aircraft [kg] mfuel - Mass of fuel [kg] minitial - Initial mass of the aircraft [kg] t - Time [s] v - Velocity [m/s; km/h] vR - Safe flying speed [m/s; km/h] vStall - Stall speed [m/s; km/h]

Appendix TABLE A1 Britten Norman BN2 data from the manufacturer

Crew

1

Passengers

9

Propulsion

2 Piston Engines

Engine Model

Lycoming O-540-E4C5

Engine Power (each)

194 Kw

260 hp

Speed

272 km/h

147 kts

Service Ceiling

4.450 m

14.600 ft

Range

1.400 km

756 NM

Empty Weight

1.638 kg

3.611 lbs

max. Takeoff Weight

2.994 kg

6.600 lbs

Wing Span

14,94 m

49 ft 0 in

Wing Area

30,2 m2

325 ft2

Length

10,86 m

35 ft 8 in

Height

4,18 m

13 ft 9 in

First Flight

13.06.1965

Production Status

in production

Total Production

1280

ICAO Code

BN2P BN2T

IATA Code

BNI

FAA TCDS

A17EU

Data for (Version)

Britten-Norman BN-2B

169 mph


870 mi.




15


TABLE A2 Battery SB4850 performance table

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