Electric aircraft have generated increased interest following the recent success of electric passenger vehicles. Over 4 million passenger electric vehicles have been sold,(1) and there have been numerous announcements regarding the electrification of SUVs, pick-up trucks, and other light commercial vehicles, which represent the majority of the passenger automotive market.(2,3) However, while electrification of ground vehicles is well underway, electrification of aircraft is still in its infancy. Conventional aircraft engines emit greenhouse gases such as carbon dioxide, water vapor, nitrous oxides, sulfates, and soot.(4) They also emit contrails, which could cause up to 50% of aviation-derived radiative forcing.(5) In addition, electrification of aircraft opens new architectures for improving efficiency such as distributed electric propulsion, which can increase the lift–drag ratio and decrease the weight of the propulsion system,(6,7) and boundary layer ingestion, which can increase propulsive efficiency by 8–10%.(8,9) Alongside, there is great interest in electric vertical takeoff and landing (eVTOL) aircraft for urban air mobility.(10−12) In a recent Viewpoint, we identified the challenging battery requirements for eVTOL aircraft, reiterating the obvious importance of specific energy (defined as the energy available per unit mass) and identifying the importance of power limitations and thermal management requirements during takeoff and landing.(13) While eVTOLs represent a new market for electric aircraft, electrifying existing commercial aircraft is an important step in moving the transportation sector toward net-zero emissions.(14) Efforts are underway toward the introduction and development of electric and hybrid electric commercial aircraft.(15−17) Norway has announced its intention to electrify its entire fleet of aircraft in the near future.(18) Further, the world’s largest seaplane operator, Harbor Air, announced their intention to electrify their fleet.(19) Numerous technological challenges remain in the electrification of aircraft, one of the primary uncertainties being the performance metrics required of batteries to do so. Many analyses have presented a comprehensive system-level perspective on transport-sized hybrid and electric aircraft and have identified subsystem component targets for the systems that they analyze.(20−23) Others have presented analyses on greenhouse gas emission reductions, resulting from electrification of aircraft.(24,25) Additionally, some small electric aircraft exist in various stages of the development process.(11,26,27) These analyses tend to be specific to certain classes of transport aircraft, and sometimes specific aircraft, rather than addressing the commercial aviation market as a whole. In this Viewpoint, we aim to identify a comprehensive set of performance metrics required for next-generation batteries to electrify commercial aircraft. We divide commercial aircraft into three categories: regional, narrow-body, and wide-body. Regional aircraft typically fly short missions, about 500 nautical miles (nmi) and carry low passenger loads (30–75), while wide-body aircraft carry high passenger loads (200–400) and fly much longer missions (>2000 nmi). Narrow-body aircraft fall in between, carrying medium passenger loads and flying ranges of ∼1000 nmi. We find that the major factor in determining the specific energy required of aircraft is the range that the class of aircraft typically flies, meaning that smaller, short-range aircraft will require less demanding battery performance metrics than larger, longer-range aircraft. We find that only next-generation chemistries, like Li–air or Li–CF_x, may be able to meet some of the requirements needed for electric commercial aircraft to achieve the range and payloads required for adoption. In the course of a mission, an aircraft takes off from the ground, climbs to its cruising altitude, cruises to its destination, descends to near ground level, and then lands.(28) All aircraft are mandated to maintain an emergency reserve energy for contingencies such as diversions or aborted landings. The FAA (Federal Aviation Administration) requires that commercial aircraft be able to abort a landing, climb to normal cruising altitude, fly to the most distant alternate airport (here assumed to be 200 nmi), and loiter for 45 min at normal cruise fuel consumption.(29) As an alternative to the extant FAA commercial reserve, a proposed approach is to house the emergency reserve by maintaining an additional 30% battery state of charge (SoC),(20) and we use the 30% SoC reserve in our analyses. To calculate energy and power requirements in flight, we first calculate thrust. Four forces act on an aircraft in flight: thrust (force generated by the propulsion system), drag (aerodynamic force opposite to velocity), weight (gravitational force), and lift (aerodynamic force normal to velocity).(30) Neglecting acceleration, thrust can be calculated by solving the equations for each of these forces, which are a function of the geometry and operating conditions of the aircraft, including instantaneous velocity relative to the surrounding air (V), the zero lift drag coefficient (C_D0), propulsive efficiency (η_prop), mechanical efficiency (η_mech), wing area (S), the aspect ratio (the ratio of the square of the wingspan to the wing area), and climb or descent angle (γ). To calculate power at any point during flight, we neglect acceleration, and thrust is multiplied by velocity, resulting in eq 1.(31) We calculate energy by integrating instantaneous power over the duration of the flight.(1) Assessing the performance of potential electric aircraft is complicated by the considerable variation of the parameters in eq 1. We calculate density (ρ) and velocity (V) at each point during the flight. We estimate the remaining parameters, namely, the zero-lift drag coefficient (C_D0), propulsive efficiency(η_prop), mechanical efficiency (η_mech), wing area (S), and aspect ratio, using distributions based on current commercial aircraft. K is a function of aircraft geometry and is discussed in more detail in the SI. To estimate mass allocated to payload, energy storage, and aircraft systems, we use the empty mass fraction (ewf), the fraction of aircraft mass with no payload or energy storage to the total takeoff mass of the aircraft. The total mass of the aircraft M_TO is given by eq 2, where M_pax is the payload mass, S_e is specific energy, and P is the instantaneous power. Detailed descriptions of the calculations of each parameter are available in the SI.(2) Figure 1. Histograms of specific energy for regional, narrow-body, and wide-body aircraft, illustrating the uncertainty stemming from aircraft design parameters. Larger (and longer-range) aircraft require a higher specific energy than do smaller (and shorter-range) aircraft. Figure 2 shows the distributions of parameters from historical aircraft showing minimum, maximum, and mean values for each parameter and class of aircraft. These parameters were gathered from current U.S. commercial aircraft (a specific list of aircraft can be found in the SI) and were used in lieu of extensive trade studies to estimate the parameters of potential electric aircraft. Figure 2. Parameters used to estimate specific energy for various classes of commercial aircraft. The minimum, maximum, and mean of each parameter and aircraft are shown on each plot. These parameters are used to estimate the power and energy of a prospective electric aircraft of each size. The data for these parameters are in the SI. To find the distribution of specific energy resulting from the uncertain parameters in eq 1, we performed Monte Carlo simulations with predefined missions for each class of aircraft. The parameters for the simulations were sampled from the triangular distributions shown in (Figure 2). The range for regional, narrow-body, and wide-body aircraft was set at 350, 500, and 1000 nmi, respectively, and the number of passengers was set to 30, 150, and 300, while the mass for each segment was 50 000, 100 000, and 250,000 kg, chosen based on previous literature and current aircraft of each class. Then, 100 000 iterations were run for each class of aircraft. Results are shown in Figure 1. The data shown in Figure 1 have means for each segment of aircraft of ∼600, 820, and 1280 Wh/kg-pack, with standard deviations of 61, 81, and 105 Wh/kg-pack, respectively. Gnadt et. al estimated a required specific energy for a narrow-body aircraft of 800 Wh/kg-pack, which agrees well with our mean for that class of 820 Wh/kg.(20) The trend of increasing specific energy with aircraft size is not primarily due to the larger size of these aircraft but rather to the longer-range use cases for which they are typically employed. When the range for the narrow-body and wide-body cases is held constant and the same analysis is run, the mean specific energy for the wide-body is ∼1280 Wh/kg-pack, and that for the narrow-body is ∼1490 Wh/kg-pack. The resulting histograms can be seen in the SI. It should be noted that satisfying these predefined mission requirements does not guarantee that an aircraft is commercially feasible. Small aircraft, such as regional and some narrow-body aircraft, often have cruising ranges that are much lower than their maximum range. However, large aircraft often use a much larger fraction of their maximum range in a typical flight. For example, the Airbus A319, a small narrow-body aircraft, is most likely to fly a range of around 161 nmi in cruise, and the distribution of its flights is skewed toward the lower end of its range. On the other hand, a wide-body aircraft, such as the Boeing 777, nearly always flies toward the high end of its range, with a mode cruise length of 2615 nmi.(32) Therefore, not only are the battery requirements for regional aircraft more feasible than narrow- or wide-body aircraft but the baseline cases for regional aircraft are also more practical than those for narrow- and wide-body aircraft. To compare potential electric aircraft and conventional aircraft at various battery-specific energies and empty weight fractions, we show the percentage of mean range and passenger nautical miles (pnmi) for each class of aircraft in Figure 3. The regional aircraft is able to achieve the current mean pnmi at around 1400 Wh/kg-pack, with an empty weight fraction of 0.35, while the narrow-body and wide-body aircraft are not able to achieve the current mean pnmi at any specific energy considered in this analysis. The most demanding battery requirements occur in the wide-body case, where even in the most optimistic case presented in this paper only 24% of the current pnmi and 20% of the current range are achieved. Figure 3. Range and passenger miles achieved by electric regional, narrow-body, and wide-body aircraft shown as a fraction of the current average range in (a) and passenger miles in (b) for the respective categories. We observe that for regional aircraft the current average range is achieved at a pack-level specific energy of about 2000 Wh/kg and current average passenger miles at about 1400 Wh/kg. The threshold for a feasible all-electric regional aircraft is about 500 Wh/kg, achieving about 25% of the current average range. A similar threshold is about 800 and 1700 Wh/kg for narrow-body and wide-body, respectively. However, only 12 and 16% of the current average range is achieved at threshold-specific energies for the narrow- and wide-body aircraft, respectively. At the highest pack-level specific energy considered of 2000 Wh/kg, electric wide-body aircraft can achieve only 19 and 16% of the current average range and passenger miles. On the other hand, at 2000 Wh/kg, regional aircraft achieve a much higher range and passenger miles than the current average. As mentioned above, the scaling effect is not primarily due to the larger size of the aircraft but rather due to the increased range. To illustrate this effect, consider the power profiles of each of the classes of aircraft for representative ranges flown by each (Figure 4). While the size of the aircraft results in the higher power at each point, the energy required to fly the ranges flown by aircraft (the area beneath the curve) increases as a result of both the increased power and the increased range. Figure 4. (a) Aircraft power profiles, along with conditions of flight in each segment. This figure illustrates the scaling challenges inherent in electric flight; as MTOM increases, the typical use case range also increases, causing a massive increase in the total energy needed. (b) Comparison of the power demand to energy (total energy over the trip) ratio throughout the mission. Having identified the energy and power requirements, we discuss the possible battery chemistries and materials needed to achieve the previously identified targets. The specific energy of current generation Li-ion batteries is about 250 Wh/kg-cell, which has steadily increased by about 5% over the past decade.(33) The projected maximum specific energy for future Li-ion batteries is around 400–500 Wh/kg-cell(33) with lithium metal anodes and high-voltage and high specific capacity cathodes. Accounting for packing burden, this is likely insufficient for regional aircraft, the least demanding among the three categories of aircraft considered. The maximum specific energy of a Li–S system is about 500 Wh/kg-pack,(34) which reaches the minimum threshold for regional aircraft, but does not allow for improvements beyond the baseline capability and therefore may not be practical for aircraft development. One of the most promising chemistries is Li–O₂, where the projected maximum pack specific energy could potentially meet some of the targets estimated previously for narrow-body and regional aircraft and allow for improvements beyond the baseline capability in the case of regional aircraft. While Li–O₂ battery systems have one of the highest specific energies among rechargeable electrochemical batteries,(34) comparable high specific energy primary batteries have been investigated for applications in space exploration.(35) At an operating temperature of about 20 °C, Li/SO₂, Li/SOCl₂, Li/FeS₂, and Li/MnO₂ systems provide specific energies in the range of 350–420 and 330–350 Wh/kg-cell at low and medium discharge rates, respectively. Li/CF_x batteries provide up to 730 Wh/kg-cell at medium discharge rates.(35) It remains to be seen if these primary battery chemistries could be made rechargeable and meet the power and specific energy requirements for electric aircraft. In this study, we limit our analysis only to rechargeable batteries for aircraft propulsion, and we intend to explore the performance envelope of these primary batteries in a future study. To estimate the cell and pack-level specific energy of Li–O₂ systems, we used electrochemical Li–air cell and pack models following the work of Gallagher et al.(34) Both open and closed Li–O₂ systems were considered for this analysis. We chose to focus on an open system, which does not carry oxygen on-board, as opposed to a closed system wherein the O₂ is contained in a pressure vessel because the open system tends to maximize specific energy, although oxygen intake over the course of a discharge will cause the mass of the system to rise over the duration of a flight, resulting in reduced effective specific energy.(34) In such a system, the battery is accompanied by a compressor to account for the changes in atmospheric pressure experienced by an aircraft in flight. The mass of the compressor and the energy that it consumes are accounted for in the model. Using the electrochemical and pack design model, we construct a Ragone plot showing the relationship between the pack specific energy and specific power, seen in Figure 5b. Li–O₂ is capable of providing the specific energy required for regional and many narrow-body flights; however in some cases, the high power requirements of takeoff will limit the specific energy of the battery. Figure 5. (a) Pack-level specific energy required for various aircraft configurations as a function of power/energy ratio and specific energy achieved by a Li–air battery as a function of power/energy ratio. While for low values of the power–mass ratio (C) all three aircraft could be flown using Li–air batteries, only for regional is a meaningful percentage of current passenger nautical miles achieved. (b) Pack specific energy of Li–air open systems for different pack-level energy and power metrics. As the specific energy tends to zero, it implies that the pack power to pack energy ratio is not achievable. Figure 5a shows the specific energy as a function of power–energy ratio for the Li–O₂ system. It also shows the specific energy required as a function of the peak power to energy (in W/Wh) ratio for each type of aircraft for various values of power–mass ratio (in W/kg). The intersection of these curves represents a feasible operating point for a prospective aircraft, where the battery power and energy meet the aircraft’s requirements. For all three categories of aircraft, only the lowest power–mass ratio (150) yields a feasible specific energy. For regional aircraft, the specific energy is around 900 Wh/kg-pack, meaning that a lithium air battery could achieve around 60% of the current passenger nautical miles according to Figure 3. For narrow-body aircraft, the maximum specific energy achieved is around 600 Wh/kg, achieving around 10% of the current passenger nautical miles, and for wide-body, no meaningful aircraft can be built at the specific energy identified. Therefore, Li–O₂ provides a feasible route forward only for small regional aircraft. Fully electric aircraft powered by batteries face a number of challenges moving forward. The specific energy of even the most optimistic future batteries enables only small regional aircraft, while larger narrow-body or wide-body aircraft remain outside of the feasibility limits of known electrochemical rechargeable battery systems. Additionally, the achievable small electric aircraft would be heavier than conventional aircraft for comparable performance metrics. It should be noted that this analysis does not consider the energy savings through potential improvements in aircraft design such as boundary layer ingestion and distributed propulsion. Although these technologies could be achieved in conventional aircraft, electrification provides the most feasible avenue for their introduction.(7) While a fully electric aircraft requires significant innovation in battery and aircraft design, a hybrid aircraft(36) could be a potential pathway to help address some of the challenges while increasing aircraft efficiency. In any case, a fully, or at least a more electric, (hybrid) aircraft presents an opportunity to lower the climate impact of commercial aviation. While the exact extent of emission savings depends on external factors such as electricity mix, electrifying aircraft would eliminate aircraft-induced cloudiness. In the near term, hybrid electric and small fully electric aircraft can help mitigate these climatic effects of aviation. In the long term, significant technological improvements in both battery and aircraft technology will aid the further adoption of small electric and larger more electric aircraft. The Supporting Information is available free of charge at https://pubs.acs.org/doi/10.1021/acsenergylett.9b02574.

• Details of the aircraft performance model, the PNMi optimization, and a comparison of narrow-body and regional aircraft specific energies at constant range (PDF)
• Aircraft parameters (TXT)

Details of the aircraft performance model, the PNMi optimization, and a comparison of narrow-body and regional aircraft specific energies at constant range (PDF) Aircraft parameters (TXT) 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): Venkat Viswanathan is a consultant for Pratt & Whitney. He is a technical consultant, owns stock options, and is a member of the Advisory Board at Zunum Aero. Viswanathan's group received research funding from Airbus A3 and Aurora Flight Sciences. Electronic Supporting Information files are available without a subscription to ACS Web Editions. The American Chemical Society holds a copyright ownership interest in any copyrightable Supporting Information. Files available from the ACS website may be downloaded for personal use only. Users are not otherwise permitted to reproduce, republish, redistribute, or sell any Supporting Information from the ACS website, either in whole or in part, in either machine-readable form or any other form without permission from the American Chemical Society. For permission to reproduce, republish and redistribute this material, requesters must process their own requests via the RightsLink permission system. Information about how to use the RightsLink permission system can be found at http://pubs.acs.org/page/copyright/permissions.html. A.B. and V.V. gratefully acknowledge support of part of this work from Convergent Aeronautics Solutions (CAS) project under the NASA Aeronautics Research Mission Directorate. This article references 36 other publications.

中文翻译：

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下一代电池使商用飞机电气化所需的性能指标

随着电动乘用车的最近成功，电动飞机引起了越来越多的兴趣。售出了超过400万辆乘用车，（1），关于SUV，皮卡车和其他轻型商用车电动化的公告很多，这代表了乘用车市场的大部分。（2,3 ）但是，尽管地面车辆的电气化进展良好，但飞机的电气化仍处于起步阶段。常规的飞机发动机排放诸如二氧化碳，水蒸气，一氧化二氮，硫酸盐和烟灰之类的温室气体。（4）它们还排放凝结尾迹，这可能导致高达50％的航空辐射强迫。（5）此外，飞机电气化为提高效率开辟了新的架构，例如分布式电力推进，（14）正在努力引进和开发电动和混合动力商用飞机。（15-17）挪威宣布打算在不久的将来使整个飞机机队电气化。（18）此外，世界上最大的水上飞机运营商Harbor Air宣布了其机队电气化的意图。（19）飞机电气化方面仍存在许多技术挑战，其中主要的不确定因素之一是电池所需的性能指标。许多分析提出了关于运输规模混合动力和电动飞机的系统级综合观点，并确定了所分析系统的子系统组件目标。（20-23）其他分析则提出了因电动化而产生的温室气体减排量的分析。飞机。（24,25）另外，在发展过程的各个阶段都存在一些小型电动飞机。（11,26,27）这些分析往往针对特定类别的运输机，有时甚至针对特定飞机，而不是针对整个商业航空市场。在此观点中，我们旨在确定下一代电池使民航飞机电气化所需的一套全面的性能指标。我们将商用飞机分为三类：支线飞机，窄体飞机和宽体飞机。支线飞机通常执行短任务，大约500海里（nmi），载客量低（30–75），而宽体飞机则载客量高（200–400），飞行时间长（> 2000 nmi）。窄体飞机介于两者之间，载有中等乘客负载，飞行范围约为1000 nmi。我们发现，确定飞机所需比能量的主要因素是该类飞机通常飞行的航程，这意味着较小的，短程飞机比大型，长程飞机需要的电池性能指标要少。我们发现只有Li-air或Li-CF等下一代化学物质_X，也许能够满足电动商用飞机达到采用所需要的航程和有效载荷所需的一些要求。在执行任务的过程中，一架飞机从地面起飞，爬升到巡航高度，巡航到目的地，下降到接近地面的高度，然后着陆。（28）所有飞机都被要求为飞机保持紧急备用能量紧急情况，例如改道或登陆失败。美国联邦航空局（FAA）要求商用飞机必须中止降落，爬升至正常巡航高度，飞往最远的候补机场（此处假定为200海里），并在正常巡航燃油消耗下游荡45分钟。（29）作为现有FAA商业储备的替代方案，一种建议的方法是通过维持额外的30％电池充电状态（SoC）来容纳应急储备，（20），我们在分析中使用30％SoC储备。为了计算飞行中的能量和功率需求，我们首先计算推力。四个力作用于飞行中的飞机：推力（由推进系统产生的力），阻力（与速度相反的空气动力），重量（重力）和升力（与速度垂直的空气动力）。（30）忽略加速度，可以通过求解这些力中的每一个的方程来计算推力，这些方程是飞机几何形状和运行条件的函数，包括相对于周围空气的瞬时速度（我们首先计算推力。四个力作用于飞行中的飞机：推力（由推进系统产生的力），阻力（与速度相反的空气动力），重量（重力）和升力（与速度垂直的空气动力）。（30）忽略加速度，可以通过求解这些力中的每一个的方程来计算推力，这些方程是飞机几何形状和运行条件的函数，包括相对于周围空气的瞬时速度（我们首先计算推力。四种力作用在飞行中的飞机上：推力（由推进系统产生的力），阻力（与速度相反的空气动力），重量（重力）和升力（与速度垂直的空气动力）。（30）忽略加速度，可以通过求解这些力中的每一个的方程来计算推力，这些方程是飞机几何形状和运行条件的函数，包括相对于周围空气的瞬时速度（V），零升阻系数（C _D0），推进效率（_ηprop），机械效率（_ηmech），机翼面积（S），纵横比（机翼面积与机翼面积的平方之比） ，以及上升或下降角度（γ）。为了计算飞行过程中任何一点的功率，我们忽略加速度，而将推力乘以速度，得出等式1。（31）我们通过在飞行过程中积分瞬时功率来计算能量。（1） 方程1中参数的显着变化使评估潜在电动飞机的性能变得复杂。我们在飞行过程中的每个点计算密度（ρ）和速度（V）。我们使用基于当前商用飞机的分布估算剩余参数，即零升阻系数（C _D0），推进效率（_ηprop），机械效率（_ηmech），机翼面积（S）和长宽比。 。ķ是飞机几何形状的函数，将在SI中进行更详细的讨论。为了估算分配给有效载荷，能量存储和飞机系统的质量，我们使用空质量分数（ewf），即没有有效载荷或能量存储的飞机质量占飞机总起飞质量的比例。飞机的总质量M _TO由等式2给出，其中M _pax是有效载荷质量，S _e是比能量，P是瞬时功率。SI中提供了每个参数计算的详细说明。（2）图1.支线飞机，窄体飞机和宽体飞机的比能量直方图，说明了飞机设计参数带来的不确定性。大型（远程）飞机比小型（远程）飞机需要更高的比能量。图2显示了历史飞机的参数分布，其中显示了每种参数和飞机类别的最小值，最大值和平均值。这些参数是从当前的美国商用飞机（在SI中可以找到特定的飞机列表）中收集的，并用于代替广泛的贸易研究来估计潜在的电动飞机的参数。图2.用于估算各类商用飞机的比能量的参数。每个参数和飞行器的最小，最大和平均值显示在每个图上。这些参数用于估计每种尺寸的预期电动飞机的功率和能量。这些参数的数据在SI中。为了找到由等式1中的不确定参数导致的比能量分布，我们对每种类型的飞机进行了具有预定义任务的蒙特卡洛模拟。仿真参数是从如图2所示的三角形分布中取样的。支线飞机，窄体飞机和宽体飞机的航程分别设置为350、500和1000 nmi，乘客人数分别设置为30、150和300 nmi，每段的质量为50 000、100 000和250,000千克，是根据以前的文献和当前每类飞机选择的。然后，为每架飞机运行了10万次迭代。结果如图1所示。图1所示数据对飞机的每个航段的均值约为600、820和1280 Wh / kg-pack，标准偏差分别为61、81和105 Wh / kg-pack。纳特（Gnadt）等 等人估计了800 Wh / kg-pack的窄体飞机所需的比能量，这与我们对于820 Wh / kg的此类平均值的平均值非常吻合。（20）随着飞机尺寸的增加，比能量的增长趋势并非主要由于这些飞机的尺寸较大，而是由于通常使用它们的航程较远。当窄体和宽体情况的范围保持恒定并进行相同的分析时，宽体的平均比能为〜1280 Wh / kg-pack，而窄体的平均比能为〜 1490 Wh / kg-pack。生成的直方图可以在SI中看到。应当指出，满足这些预定的任务要求并不能保证飞机在商业上是可行的。小型飞机（例如支线飞机和某些窄体飞机）的巡航距离通常远低于其最大飞行距离。但是，大型飞机在典型飞行中通常会使用最大范围的很大一部分。例如，空中客车A319是一种小型窄体飞机，最有可能在巡航时飞行约161 nmi的航程，并且其飞行分布偏向其航程的下限。另一方面，像波音777这样的宽体飞机几乎总是向其航程的高端飞行，其模式巡航长度为2615 nmi。（32）因此，不仅支线飞机的电池需求比窄体或宽体飞机更可行，而且支线飞机的基线情况也比窄体和宽体飞机的实际情况更为实用。为了比较在各种电池特定的能量和空载重量分数下的潜在电动飞机和常规飞机，我们在图3中显示了每种飞机的平均航程和乘客海里（pnmi）的百分比。该地区飞机能够实现当前平均pnmi约为1400 Wh / kg-pack，空载重量分数为0.35，而窄体和宽体飞机在此分析中考虑的任何特定能量下均无法达到当前平均pnmi。电池盒最苛刻的要求是 即使在本文介绍的最乐观的情况下，也只能实现24％的电流pnmi和20％的电流范围。图3.区域性电动飞机，窄体和宽体飞机实现的航程和乘客英里数，分别表示（a）和（b）中当前平均航程的几分之一。我们观察到，对于支线飞机而言，当前的平均射程是在约2000 Wh / kg的装箱级比能量下实现的，而目前的平均乘客英里数约为1400 Wh / kg。可行的全电动支线飞机的阈值约为500 Wh / kg，达到当前平均航程的25％。窄体和宽体的相似阈值分别约为800和1700 Wh / kg。然而，对于窄体飞机和宽体飞机，分别达到阈值特定的能量时，分别只能达到当前平均范围的12％和16％。考虑到最高的包装水平比能量为2000 Wh / kg，电动宽体飞机只能达到当前平均航程和乘客里程的19％和16％。另一方面，支线飞机在2000 Wh / kg的情况下，航程和乘客英里数比目前的平均水平高得多。如上所述，缩放效果主要不是由于飞机尺寸较大，而是由于航程增加。为了说明这种效果，请考虑每种飞机所代表的航程的每种飞机的功率曲线（图4）。飞机的尺寸会导致每个点的功率更高，由于功率增加和航程增加，飞行飞机飞行的航程（曲线下方的区域）所需的能量也会增加。图4.（a）飞机功率曲线以及每个航段的飞行条件。该图说明了电动飞行固有的缩放挑战；随着MTOM的增加，典型的用例范围也会增加，从而导致所需的总能量大量增加。（b）整个任务期间的电力需求与能量（行程总能量）之比。确定了能量和功率要求后，我们讨论了实现先前确定的目标所需的可能的电池化学和材料。当前一代锂离子电池的比能量约为250 Wh / kg-电池，在过去十年中稳定增长了5％。（33）未来锂离子电池的最大比能量预计约为400-500 Wh / kg-cell（33），带有锂金属阳极，高压和高比容量阴极。考虑到包装负担，这对于支线飞机而言可能是不够的，这在所考虑的三类飞机中要求最少。Li-S系统的最大比能约为500 Wh / kg-pack，（34）达到了支线飞机的最低阈值，但不允许进行超出基准能力的改进，因此对于飞机的开发可能不切实际。Li–O是最有前途的化学之一 对于支线飞机而言，这可能是不够的，这在所考虑的三类飞机中要求最少。Li-S系统的最大比能约为500 Wh / kg-pack，（34）达到了支线飞机的最小阈值，但不允许超出基准能力进行改进，因此对于飞机的开发可能不切实际。Li–O是最有前途的化学之一 对于支线飞机而言，这可能是不够的，这在所考虑的三类飞机中要求最少。Li-S系统的最大比能约为500 Wh / kg-pack，（34）达到了支线飞机的最低阈值，但不允许进行超出基准能力的改进，因此对于飞机的开发可能不切实际。Li–O是最有前途的化学之一_{如图2所示}，在这种情况下，预计的最大装箱比能量可能会达到先前为窄体和支线飞机估计的一些目标，并且在支线飞机的情况下可以进行超出基准能力的改进。尽管Li–O ₂电池系统是可再充电电化学电池中最高的比能量之一，但[34]已研究了可比的高比能量一次电池用于太空探索。[35]在约20°C的工作温度下， Li / SO ₂，Li / SOCl ₂，Li / FeS ₂和Li / MnO ₂系统在低和中放电速率下分别提供350-420和330-350 Wh / kg电池的比能。锂/ CF_x电池在中等放电速率下可提供高达730 Wh / kg的电池电量。（35）这些初级电池化学物质能否制成可充电电池并满足电动飞机的功率和特定能量要求，还有待观察。在这项研究中，我们的分析仅限于飞机推进用的可充电电池，并且我们打算在以后的研究中探索这些一次电池的性能范围。为了估计Li–O ₂系统的电池和电池组比能，我们根据Gallagher等人的工作[34]使用电化学Li–空气电池和电池组模型。[34]开放式和封闭式Li–O ₂系统都考虑了这个分析。我们选择将重点放在开放系统上，该系统不携带氧气，而封闭系统中的O尽管打开过程中的氧气摄入会导致系统质量在飞行过程中上升，但开放式系统会趋于使比能最大化，因此将图₂包含在压力容器中。（ 34）在这种系统中，电池配有压缩机，以说明飞行中的飞机所经历的大气压力变化。在模型中考虑了压缩机的质量及其消耗的能量。使用电化学和电池组设计模型，我们绘制了一个Ragone图，显示了电池组比能量和比功率之间的关系，如图5b所示。Li–O ₂能够提供区域性和许多窄体飞行所需的特定能量；但是，在某些情况下，起飞的高功率要求将限制电池的比能量。图5.（a）各种飞机配置所需的装箱级比能量与功率/能量比的函数关系，以及锂空气电池获得的比能量与功率/能量比的函数关系。而对于低的功率质量比（C）所有这三架飞机都可以使用锂空气电池飞行，仅在区域范围内实现了当前乘客海里的可观百分比。（b）锂空气开放系统的电池组特定能量，用于不同的电池组级能量和功率指标。由于比能趋于零，这意味着无法达到电池组功率与电池组能量之比。图5a显示了Li–O _2的比能量与功率-能量比的关系系统。它还显示了在各种功率质量比值（以W / kg为单位）下，每种类型飞机所需的特定能量随峰值功率与能量（以W / Wh为单位）之比的函数。这些曲线的交点代表了预期飞机的可行工作点，其中电池功率和能量满足了飞机的要求。对于所有这三类飞机，只有最低的功率质量比（150）才能产生可​​行的比能量。对于支线飞机，比能量约为900 Wh / kg-pack，这意味着根据图3，锂空气电池可以达到当前乘客海里的60％左右。对于窄体飞机，最大比能量为约600 Wh / kg，达到目前乘客海里的10％，对于宽体来说，以确定的特定能量无法制造出有意义的飞机。因此，Li–O₂提供仅适用于小型支线飞机的可行路线。由电池供电的全电动飞机在前进中面临许多挑战。即使是最乐观的未来电池，其比能量也仅能使小型支线飞机使用，而较大的窄体或宽体飞机则不在已知的电化学可充电电池系统的可行性范围之内。另外，对于可比较的性能指标，可实现的小型电动飞机将比常规飞机重。应当指出的是，该分析并未考虑通过飞机设计的潜在改进（例如边界层吸入和分布式推进）而节省的能源。尽管这些技术可以在常规飞机上实现，电气化为引入它们提供了最可行的途径。（7）虽然全电动飞机需要电池和飞机设计方面的重大创新，但混合动力飞机（36）可能是帮助解决某些挑战并提高飞机效率的潜在途径。在任何情况下，一架完全或至少是电动的（混合动力）飞机都为降低民用航空对气候的影响提供了机会。虽然减排的确切程度取决于外部因素，例如电力混合，但飞机电气化将消除飞机引起的混浊。在短期内，混合动力电动飞机和小型全电动飞机可以帮助减轻航空的这些气候影响。在长期，电池和飞机技术的重大技术进步将有助于小型电动飞机和大型电动飞机的进一步采用。可从https://pubs.acs.org/doi/10.1021/acsenergylett.9b02574免费获得支持信息。

• 飞机性能模型，PNMi优化的详细信息，以及恒定范围内窄体和地区飞机特定能量的比较（PDF）
• 飞机参数（TXT）

飞机性能模型的详细信息，PNMi优化以及恒定范围内窄体飞机和支线飞机特定能量的比较（PDF）飞机参数（TXT）此观点中的观点是作者的观点，不一定代表ACS。作者声明存在以下竞标的财务利益：Venkat Viswanathan是普惠公司的顾问。他是技术顾问，拥有股票期权，并且是Zunum Aero咨询委员会的成员。Viswanathan的小组获得了空中客车A3和Aurora Flight Sciences的研究资助。无需订阅ACS Web版本即可获得电子支持信息文件。美国化学学会在任何可版权保护的支持信息中拥有版权权益。ACS网站上提供的文件只能下载供个人使用。未经美国化学学会许可，不得以其他方式允许用户以机器可读形式或任何其他形式全部或部分复制，重新发布，重新分发或出售ACS网站上的任何支持信息。为了获得复制，重新发布和重新分发此材料的许可，请求者必须通过RightsLink许可系统处理自己的请求。有关如何使用RightsLink权限系统的信息，请访问http://pubs.acs.org/page/copyright/permissions.html。AB和VV非常感谢NASA航空研究任务局下的Convergent Aeronautics Solutions（CAS）项目对这项工作的支持。本文引用了其他36个出版物。