Performance Metrics Required of Next-Generation Batteries to Electrify

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Performance Metrics Required of NextGeneration Batteries to Electrify Vertical Takeoff and Landing (VTOL) Aircraft

ACS Energy Lett. Downloaded from pubs.acs.org by 5.62.157.28 on 11/20/18. For personal use only.

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also has a set of opposing forces that it must counter with propulsive forces. In the terminology of flight dynamics, the force counteracting gravity is referred to as lift. Drag, exerted by the medium, acts in the direction opposite to motion. Rotors and propellers are essentially rotating blades that produce a net force, thrust, by displacing air. Rotors act in the vertical direction to produce lift like a conventional helicopter, while propellers act in the forward direction without generating significant lift. This means that propellers must be paired with wings, which will produce lift as the propellers move the aircraft forward. VTOL, which stands for “Vertical Takeoff and Landing”, refers to a class of aircraft that use rotors to take off and land, like the name suggests, without needing a runway. Helicopters are the most common example of this in existence today, though some winged aircraft employ VTOL strategies as well.15 Their operational flexibility makes VTOL aircraft suitable for the emerging potential for air transportation to replace some of the automobile traffic in cities.16 Advances in battery technology have led to an increasing EV market share,17 and similar efforts have translated to the advent of electric aircraft. Electrification comes with a number of advantages. The use of an electric powertrain will likely reduce operating costs through improvements in energy efficiency18 and potentially lower maintenance costs similar to EVs.19 Also, because the efficiency and power density of electric motors does not vary significantly with size, the propelling force or thrust can be spread out over several small motors on the aircraft body. This permits design layouts that can exploit favorable aerodynamic effects, which are otherwise not feasible with combustion engines.20,21 Furthermore, electric powertrains have additional benefits of lower noise22,23 and no exhaust pollution. Previous studies18,22 have identified four main VTOL configurations that can be powered by an electric energy source, namely, (i) a helicopter design that uses a rotor to provide lift throughout the mission, (ii) a “stopped-rotor” design that uses separate rotors and propellers for the VTOL and cruise segments, respectively, (iii) a “tilt-rotor” design in which the rotors used in the vertical flight segments rotate to become propellers for forward flight, and (iv) a “tilt-wing” design, which is similar to a tilt-rotor but with an entire wing− motor structure rotating rather than just the motors. The aforementioned analyses showed that helicopter designs tend to have higher energy requirements compared to the other designs for a given range. This is due to the other designs’ 20%

lectrification of passenger vehicles is well on its way, with global electric vehicle (EV) sales crossing 1.1 million annually1 and projections reaching 11 million by 2025. Numerous countries have also pledged a transition to all-electric fleets for new vehicles in the next few decades.2,3 Alongside, there is increasing interest toward electrifying other segments of transportation and aviation;4,5 however, a critical challenge for enabling electrification of these markets is the batteries. In a Viewpoint last year, we analyzed the battery requirements associated with electrifying heavy-duty trucks and highlighted the importance of improving the specific energy of Li-ion batteries.6 This analysis was widely covered and provides a rational basis for making engineering, design, and policy decisions. As the world’s population is increasingly moving toward cities,7 there is greater interest toward new modes of urban mobility to alleviate congestion and decrease commute times. Micromobility solutions around electric bikes and scooters are gaining traction to address the congestion challenge, alongside a growing interest toward urban air mobility. One vision of urban air mobility, as outlined by NASA, is to develop a safe and efficient system for air passenger and cargo transportation within an urban area.8 Among the many options considered for urban air mobility, electric vertical takeoff and landing (eVTOL) aircraft have emerged as near-term viable candidates with numerous efforts to commercialize such aircraft.9−14 Following a “first-principles” approach similar to our analysis of electric semitrucks, in this Viewpoint, we identify targets required of next-generation batteries (see Figure 1) to enable practical e-VTOLs. Any aircraft for flight needs to take off, climb to its flight altitude, and cruise toward its destination where it will finally land. Similar to an automobile that generates a forward force to overcome inertia, road friction, and air resistance, an aircraft

Received: November 13, 2018 Accepted: November 15, 2018

Figure 1. Pictorial illustration of the trajectory for battery pack requirements for different modes of electric mobility. © XXXX American Chemical Society

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DOI: 10.1021/acsenergylett.8b02195 ACS Energy Lett. 2018, 3, 2989−2994

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Cite This: ACS Energy Lett. 2018, 3, 2989−2994

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ACS Energy Letters decrease18 in power required for the longest segment, cruise, which results from both higher lift efficiency from wings and the lack of a tail rotor.18,24 Improvements in battery energy capabilities will extend feasible design ranges; therefore, optimizing cruise performance rather than vertical flight performance will likely continue to be more effective at reducing the overall energy use. For this reason, a generic vertical-to-fixed-wing transitioning aircraft is considered for this investigation. The details of the canonical mission considered for this analysis can be found in the Supporting Information and is informed by previous e-VTOL analyses.18,22,25 The mission comprises vertical climb, vertical descent, fixed-wing climb, fixed-wing cruise, and fixed-wing descent. In addition, aviation standards require aircraft to carry additional energy on-board as “reserve” for unforeseen events. While the Federal Aviation Administration (FAA) mandates these requirements, in the case of urban e-VTOLs, the FAA has yet to specify the concerned standards. Hence, we assume a reserve cruise amounting to 30 min based on current Instrument Flight Rules helicopter standards.26 The effect of assuming the Visual Flight Rules requirement of 20 min is discussed in the Supporting Information. Cruising range is assumed to include the fixedwing climb, cruise, and descent, where the cruise altitude is set to 1000 ft.18 Additionally, it is assumed that the aircraft will transition between vertical and fixed-wing flight 50 ft above the ground. To estimate the power requirements of the e-VTOL over the canonical mission, we employ momentum theory and standard equations of motion, corrected as necessary for electromechanical efficiency (ηmech) and propeller efficiency (ηprop), as listed in Table 1. A representative power profile is shown in

Supporting Information. The density of air (ρ) is calculated at flight altitude,28 with ground-level conditions considered between 0.974 kg/m3 as the extreme poor condition24 and 1.225 kg/m3 as the baseline.24 Disk loading (W/A) is the ratio of weight to total rotor area and is significant for the maximum power. We assume a baseline of 50 kg/m2 and varied between 35 (ideal) and 65 kg/m2 (poor). Considering fixed-wing flight next, the required power for climb, cruise, and descent can be defined as a function of the lift-to-drag ratio (L/D). This term encompasses the drag penalty for the lift needed to support the aircraft weight. While the ratio is highly sensitive to the specific structural design, there is a paucity of real-world data sets; hence, it is fixed at an estimated value of 14, which is representative of a tilt-rotor.29 The lift-to-drag ratio is related to power by eq 224 ÄÅ É ÅÅ WV ÑÑÑÑ Å Pfixed‐wing = ÅÅWVv + Ñ/(η η ) ÅÅÇ L /D ÑÑÑÖ mech prop (2) where the vertical velocity (Vv) and forward velocity (V) are segment-specific.24 As defined by the mission, the vertical velocity component (Vv) is zero for cruise, 900 fpm for climb, and −500 fpm for descent. Forward velocity is set to the minimum power velocity (VMinPower) for climb and descent.24 For cruise, velocity is set to the maximum range velocity (VMaxRange). These values depend on the wing loading (W/S), ranging from 63 to 103 kg/m2,24 and the zero-lift drag coefficient (CD0), assumed to be 0.03. These velocities are derived in the Supporting Information. L/D is also reduced by 13% during climb and descent corresponding to minimum power conditions.24 Power for both transition segments is assumed to vary linearly between the vertical and fixed-wing segments before and after, over the specified transition time of 120 s.22 The effects of alternatively assuming constant hover power for transition are discussed in the Supporting Information. Each of the parameters in Table 1 have a suitable range of values for an e-VTOL design. The values identified here as “ideal” in Figure 2 represent all of these parameters being at the most favorable value with respect to energy requirements. It should be noted that this refers to ideality within our design space and not to 100% efficiency. In actual application, these parameters interact and potentially conflict in ways that are highly dependent on structure and configuration, and this would be closer to an upper theoretical limit under the given assumptions. Most approaches to evaluate e-VTOLs18,22,25 have generally focused on using a specified range and a corresponding estimation of takeoff mass. Such a design-based approach is useful in comparing different design concepts for a specific mission; however, the goal of the current work is to analyze the general performance metrics of e-VTOL aircraft. To determine the effects of salient design parameters on range, we fix the gross takeoff mass (GTOM) of the vehicle for a given set of input parameters and determine the maximum resulting range. The mass buildup for the model is described in the Supporting Information. As a result of the above-mentioned approach, battery weight is fixed for a given set of performance parameters. It follows that by increasing the specific energy of the pack there would be a linear increase in the achievable mission range, as shown in Figure 2a. Increasing GTOM also extends the potential cruise range as long as the increase in battery energy provides

Table 1. Extreme and Baseline Values parameter

description

2

W/A (kg/m ) disk loading W/S (kg/m2) wing loading ρground (kg/m3) air density at ground level ηmech electromechanical efficiency ηprop propeller efficiency We/W empty weight ratio fundamental constraints description SE (Wh/kg) GTOM (kg) PAX (100 kg ea.)

pack-level specific energy gross takeoff mass number of passengers

[poor, ideal], baseline [65, 35], 50 [103, 63], 83 [0.974, 1.225], 1.225 [0.85, 0.95], 0.90 [0.7, 0.9], 0.8 [0.65, 0.55], 0.60 [bounds], baseline [150, 300], 150 [1000, 2500], 1000 [1,6], 1

Figure S5. For the vertical mission segments, eq 1 provides the power (P) required to lift the gross weight (W) at a given vertical climb rate (Vclimb,v).24 ÄÅ É ÅÅ fW WVclimb,v ÑÑÑÑ fW /A Å Å ÑÑ/η + Pvertical = ÅÅ ÅÅ FoM ÑÑÑ mech ρ 2 2 (1) ÅÇ ÑÖ This climb rate is 500 ft/min (fpm) for vertical climb and descent, although it is held at zero while calculating the power for vertical descent. The 0 fpm during descent provides reasonable estimates for the minimum power necessary for safe descent as the rotors move into their own “down-wash”,27 as the aircraft tries to push up off of air that has already been accelerated downward by the rotors. The figure of merit (FoM), set to 0.7,24 and fuselage down-wash correction (f), set to 1.03,24 are efficiency terms, described in detail in the 2990

DOI: 10.1021/acsenergylett.8b02195 ACS Energy Lett. 2018, 3, 2989−2994

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ACS Energy Letters

greater than the 43 mile sizing mission range that NASA estimates.25 However, baseline assumptions offer only 20 miles, while universally “poor” performance parameters under our assumptions could not produce a feasible aircraft. Increasing specific energy to 300 Wh/kg not only increases the limiting and baseline ranges to 200 and 120 miles, respectively, for the 1000 kg e-VTOL but also allows for a viable, 35 mile range aircraft even at the pessimistic end of the parameter space. Generally, more favorable design and performance would lead to a feasible e-VTOL at lower specific energies, while higher specific energy could accommodate poorer performance. Further computational modeling and testing of these novel structural and propulsive designs would narrow down the spread for defining vehicle parameters and feasibility regions, eventually leading to a better quantitative description of the energy and power required from the battery pack. The aircraft modeling provides estimates for the total required battery pack energy and the threshold specific energy. Besides these constraints, the energy available from battery packs based on Li-ion cells would have a discharge rate limitation as well.30 Furthermore, as shown in Figure 3b, not only does the takeoff segment require a high power output from the battery pack, the landing segment also requires an

Figure 2. (a) For a single-passenger e-VTOL, the cruise range increases linearly by improving the specific energy. The baseline conditions for several GTOM values within the likely spread for one passenger are given by the solid lines, while the extreme ideal and poor performance conditions represented in Table 1 are shown by the dotted and dashed lines, respectively. (b) With an SE value of 150 Wh/kg, which represents the current state-of-the-art, accommodating more people implies either dropping the maximum range or designing a larger aircraft. This chart shows the maximum length of an e-VTOL trip for an increasing number of passengers over several GTOM values. The chart focuses specifically on the ideal limit, corresponding to the dotted lines in part (a). Increasing GTOM sees diminishing returns in the range as the mass penalty of adding more batteries begins to outweigh the increase in energy.

more than sufficient energy to support the higher weight. The dotted lines in Figure 2a are an upper limit set by considering all of the ideal parameters in Table 1, the solid lines represent the baseline assumptions, and the dashed lines represent all parameters set to poor performance. As presented above, this approach is not meant to be predictive because trade-offs would make the extremes highly unlikely in real-world application. Figure 2b shows the results of fixing the packlevel specific energy at 150 Wh/kg and adjusting the mass constraints. For increasing takeoff masses, the maximum cruise range at the ideal limit is shown for an increasing number of passengers (PAX) considered in increments of 100 kg. The resulting payload-range trade-off comes from exchanging battery mass for passenger and luggage mass. The results shown in Figure 2 suggest that the feasibility of an e-VTOL is highly dependent on both specific energy and the parameters in Table 1. At a GTOM of 1000 kg and SE of 150 Wh/kg, an operational range of up to 73 miles could be achieved under conditions of ideality, which is significantly

Figure 3. Discharge performance of the cells and the battery pack simulated for new cells. (a) Discharge C-rate for the cells over the 73 mile mission, where we observe a discharge rate of about 4C for takeoff and about 4.8C for landing. The landing discharge C-rate could increase if the reserve segment is included within the mission. (b) Comparison of the ratio of power demand (P) to the pack energy (EP) for e-VTOLs, electric automobiles, and electric semitrucks, where the need for higher power is shown. (c) Comparison of the ratio of heat generation (qP) rate to the pack size (EP) for equivalent trips of e-VTOLs and EVs, where we observe an order of magnitude difference, thereby highlighting the need for novel thermal management strategies for e-VTOL batteries. 2991

DOI: 10.1021/acsenergylett.8b02195 ACS Energy Lett. 2018, 3, 2989−2994

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battery packs will focus on increasing the specific energy and thereby increasing the total pack energy for the same pack weight.32 These improvements would not necessarily have an impact on battery packs for e-VTOLs because improvements in specific energy for current Li-ion batteries entail a reduction in the maximum discharge rates and the power performance, as shown in Figure S1. An active area of research in EV batteries is the development of cells capable of fast charging without compromising the specific energy;33 however, e-VTOL aircraft batteries would also need to be capable of high-rate discharge as highlighted by Figure 3a. In order to evaluate the thermal requirements of e-VTOLs during discharge, we compare the ratio of heat released to the pack energy for 73 mile trips of e-VTOLs and EVs (an automobile and a semitruck) where each trip begins at ambient conditions of 25 °C. The ratio of the total heat generation rate (qP) to the total pack energy (EP) for the three use cases considered is shown in Figure 3c. The heat generation rate is significantly higher for e-VTOLs, where qP/EP is about 0.6 and 0.25 for landing and takeoff, respectively, in comparison to a maximum value of 0.05 and 0.002 for the electric automobile and the semitruck. Apart from the heat generation from the cells, additional heat would be generated in other powertrain components due to high currents of the landing and takeoff segments, which can potentially be mitigated through higher voltage battery packs, notwithstanding, an important aspect to consider in this context. Furthermore, the simulations shown in Figure 3c were performed on fresh cells with no cycling and aging effects. As the performance of cells deteriorates over cycling, the thermal response would change significantly due to the increase in impedance and other effects of cell degradation.34 It follows that unique thermal management solutions might be necessary for the e-VTOL applications, unlike conventional strategies used in EVs. It is worth noting that in this analysis we do not discuss the charging segments for e-VTOL batteries because the operational models and the corresponding charging requirements are uncertain at this point.35 Another extremely important aspect of analyzing the feasibility requirements of batteries for e-VTOLs requires a closer look at the end-of-life (EOL) criterion. We have already highlighted the importance of power along with the specific energy threshold for e-VTOL batteries. A reduction in capacity and total energy of the battery pack would reduce the operational range. On the other hand, a reduction in the power capability or, in other words, a reduction in the maximum discharge C-rate of the cells would render an e-VTOL aircraft incapable of taking off, and more importantly, before it reaches the takeoff limitation, it would be incapable of landing where the C-rate demands are much higher. Similar to capacity fade, the parasitic processes also increase the impedance of the cell, translating to a power fade that occurs alongside capacity fade. 34 Conventional EVs have an EOL criterion of degradation to 80% of the initial capacity.36 In the context of e-VTOLs, it is likely that the landing power requirements would pose a power fade limitation that could determine the EOL criterion. In addition, in the context of designing battery systems for aviation applications, the systems should be capable of handling battery pack failure modes. In such a case, a subset of the total modules within the battery pack would have to provide the required power during landing; therefore, the EOL conditions for power fade would have to include the performance of the pack in the stipulated failure modes.

equivalent amount of power. Because the voltage of the pack (and the cells within) drops over discharge for Li-ion batteries, the current output required from the battery pack would be much higher during landing compared to the takeoff segment. Currently available Li-ion batteries have a maximum continuous discharge C-rate specification shown in Figure S1, where we also observe that as the design specific energy of cells increases the maximum C-rate capability decreases significantly. The maximum cell specific energy currently available is well under 300 Wh/kg, as seen in Figure S1, with a maximum Crate approaching about 3C as the design specific energy exceeds 250 Wh/kg. This trend is similar, but does not correspond to the phenomenon captured by a conventional Ragone Chart. The trend is governed largely by the cell chemistry/materials and partly by cell engineering and design.30 The representative cell chosen for this study has a cell specific energy of 245 Wh/kg, and within the modeling framework, the cell can sustain C-rates of up to 5C discharge. Other characteristics of the cell are shown in Figure S2. Upon accounting for the packing burden,6,31 the cell specific energy of 245 Wh/kg would reduce to about 150−170 Wh/kg at the pack level, where we assume a packing burden factor of 0.6− 0.7.6,31 The packing burden accounts for the packing material, interconnects, and additional weight of the inactive materials required to assemble a battery pack from individual cells. In Figure 3a, we examine the discharge C-rate for the power profile of the 73 mile, ideal conditions mission, when applied to a 52.5 kWh battery pack, including a reserve of 15.5 kWh. On the pack level, we examine different pack voltages to estimate the required current output from the battery pack for the same power profile, as shown in Figure S3. We find that a pack with a nominal voltage of 500 V would require a peak current output of ∼450 A during landing, while a pack with a nominal voltage of 650 V would require ∼335 A. It should be noted that the cells within each of these battery packs would behave in a similar manner because the total number of cells is constant. The pack discharge current would be an important factor in battery pack and powertrain design considerations. Lower pack currents would facilitate the use of lighter cables, which could influence the empty weight fraction and mitigate the efficiency losses that occur at high currents. For the purpose of this study, the analysis is conducted using the 650 V battery pack due to its lower current output requirements. We can study the battery discharge performance in Figure 3 conducted on a battery pack with new cells within the modeling framework. We estimate a discharge C-rate of close to 5C (4.8C) during the landing segment, as shown in Figure 3a, compared to about 4C during the takeoff segment, which is due to the lower state-of-charge and lower voltage during landing. This implies that the limiting mission segment is not takeoff but landing where an equivalent power is required at a low state-of-charge. It should be noted that if the reserve segment were utilized, the battery pack would be at an even lower state-of-charge, which would result in a much higher discharge C-rate during landing (>5C). It is important to identify the limiting mission segment because it informs pack design and pack sizing decisions. Figure 3b features a comparison between an e-VTOL capable of 73 miles with equivalent trips for an electric semitruck with a 1000 kWh battery pack and an automobile with a 60 kWh pack. We see that the ratio of power demand (P) to pack energy (EP) is several-fold higher for e-VTOLs. Further, improvements in EV 2992

DOI: 10.1021/acsenergylett.8b02195 ACS Energy Lett. 2018, 3, 2989−2994

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ACS Energy Letters In this Viewpoint, we have examined the power requirements of an electric VTOL governed by the gross takeoff mass and the specific energy of the battery pack. On the basis of current Li-ion batteries, with a cell specific energy of around 245 Wh/kg along with a maximum discharge capability of about 5C, we constructed a 52.5 kWh Li-ion battery pack that represents the state-of-the-art. A nominal pack voltage of 650 V was used to reduce the required battery pack discharge current. We looked into a set of “ideal” conditions that correspond to high energy conversion efficiencies and assumed lightweight materials for favorable weight properties. For the canonical eVTOL considered, we estimate a takeoff discharge rate of 4C and a landing segment discharge of close to 5C due to the lower voltage of the battery pack during landing. On the basis of the discharge rates over the mission, it is evident that the limiting mission segment is landing. This highlights the need for high discharge rate along with constraints of specific energy, unlike EVs and electric heavy-duty trucks. Further, we explored the heat generation for the battery pack over the 73 mile mission and found that the ratio of the heat generation to the pack size is over an order of magnitude higher than an equivalent trip for an electric automobile and over 2 orders of magnitude higher than an electric semitruck. This comparison provides the basis for establishing the need for novel thermal management strategies for e-VTOL aircraft. Lastly, we examined the EOL criterion for e-VTOLs, where we identified the need for examining power fade, along with battery pack failure modes, which is likely to determine the EOL criterion, unlike conventional EVs where capacity (or energy) fade dictates the EOL. In conclusion, with current Li-ion battery specific energy and discharge performance, we estimate that an operational range of 73−100 miles represents the upper limit under the listed assumptions and a GTOM of 1000−2500 kg. As the specific energy of Li-ion batteries increases along with higher discharge rates, longer cruise range can be achieved for a similar GTOM; however, novel approaches are required for thermal management and EOL criteria for e-VTOL aircraft.

Venkatasubramanian Viswanathan: 0000-0003-1060-5495 Author Contributions ¶

Notes

Any views expressed here are those of the authors and do not reflect the views of Airbus, Zunum Aero, and Pratt and Whitney. Views expressed in this Viewpoint are those of the authors and not necessarily the views of the ACS. The authors declare the following competing financial interest(s): Geoffrey Bower is the Chief Engineer for Vahana at A^3 by Airbus and holds Airbus stock. Venkat Viswanathan is a consultant for Pratt & Whitney. He is a consultant, owns stock options, and is a member of the Advisory Board at Zunum Aero.

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ACKNOWLEDGMENTS W.L.F. and V.V. acknowledge support from Airbus A^3.

†

Department of Mechanical Engineering, Carnegie Mellon University, Pittsburgh, Pennsylvania 15213, United States ‡ A^3 by Airbus, Santa Clara, California 95050, United States

ASSOCIATED CONTENT

S Supporting Information *

The Supporting Information is available free of charge on the ACS Publications website at DOI: 10.1021/acsenergylett.8b02195. Details of the power consumption modeling of e-VTOL aircraft and details of the battery modeling undertaken for the study (PDF)

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REFERENCES

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William L. Fredericks†,¶ Shashank Sripad†,¶ Geoffrey C. Bower‡ Venkatasubramanian Viswanathan*,†

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W.L.F. and S.S. contributed equally to this work.

AUTHOR INFORMATION

Corresponding Author

*E-mail: [email protected]. ORCID

William L. Fredericks: 0000-0002-5069-4915 Shashank Sripad: 0000-0003-1785-2042 2993

DOI: 10.1021/acsenergylett.8b02195 ACS Energy Lett. 2018, 3, 2989−2994

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DOI: 10.1021/acsenergylett.8b02195 ACS Energy Lett. 2018, 3, 2989−2994