Electric aircraft design principles

Abstract

The present disclosure relates to an aircraft comprising a fuselage and at least one wing arranged in a wing-and-body design, a plurality of propulsors, and a primary energy source configured to power the propulsors, the primary energy source comprising a rechargeable battery. The aircraft has a maximum takeoff mass (MTOM) of at least 8,618 kg and an electric range factor (ERF) of at least 6.

Description

[Technical Field]

[0001] The present disclosure relates to design principles for an electric aircraft and a wing (and related wing assembly) for the electric aircraft. The present disclosure also relates to an electric aircraft based on the design principles and a wing (and related wing assembly) for the electric aircraft. The present disclosure also relates to an alternative backup energy source for the electric aircraft. The present disclosure also relates to a load-bearing hatch for a wing (and related wing assembly) for the electric aircraft. In particular, the electric aircraft is a wing-and-body design electric aircraft including a rechargeable battery. [Background technology]

[0002] Aviation accounts for approximately 2-3% of total carbon dioxide emissions. Aviation is expected to grow by 4-5% per year over the next decade. In addition to carbon dioxide, aviation also contributes to the creation of contrails and higher atmospheric levels of NO. X The challenge of net zero is two-fold: both carbon and non-carbon emissions must be reduced to zero.

[0003] Current flight route patterns clearly indicate which types of aircraft and distances fly most to CO2 and associated non-CO2 emissions. From ICCT estimates of CO2 emissions from commercial aviation in 2013, 2018 and 2019, it can be estimated that approximately 7% of all civil aviation CO2 emissions are attributable to flights shorter than 500 km, 19% to flights up to 1000 km and 43% to flights up to 2000 km.

[0004] Much recent literature has highlighted the perceived limitations of battery-electric aircraft.

[0005] In "Performance analysis of an electrically assisted propulsion system for a short-range civil aircraft," J Aerospace Engineering 2019, Vol. 233(4) 1490-1502, Ang et al. conclude that for short-range civil aircraft, the limited power-to-weight ratio of the electrical components hinders the development of all-electric propulsion. Instead of using all-electric propulsion, the authors argue that it would be more beneficial to use an electric system as a secondary system that operates in parallel and assists the existing primary propulsion system during certain flight phases to increase overall efficiency.

[0006] Hall et al., in "Feasibility of Electrified Propulsion for Ultra-Efficient Commercial Aircraft Final Report, NASA, 2019," conclude that not all electric designs are feasible at all scales (short to long range) because the required battery mass is greater than the aircraft parameters can support. The paper argues that even with optimistic estimates of battery capacity available in 2035, it is unlikely that currently designed missions for short-range, regional medium-range, and long-range aircraft will be powered solely by batteries.

[0007] Epstein and O'Flarity conclude in "Considerations for Reducing Aviation's CO2 with Aircraft Electric Propulsion," Journal of Propulsion and Power, May 2019, Volume 35, Number 3, that electric propulsion is not a promising path to significantly reducing aviation CO2 emissions in the first half of the 21st century. In support of this conclusion, the authors note that 92% of aviation CO2 is generated by single-aisle and twin-aisle aircraft, which require 15,000 to 200,000 kW of shaft power and 150,000 to 2,250,000 kWh of energy at takeoff, and that known battery technology is incapable of powering such aircraft for the distances currently flown.

[0008] Schafer et al., in "Technological, economic, and environmental prospects of all-electric aircraft," Nature Energy, 2018, 4, pp. 160-166, predict that electric aircraft with 10-30% battery weight will be feasible for regional or short-haul flights only if the specific energy of the batteries increases to values ​​much greater than those currently available.

[0009] In "Realizing Zero-Carbon Emission Flight—Primary Energy Source Comparison and Selection," published by the Aerospace Technology Institute, September 2021, Webber and Job argue that purely battery-powered aircraft are only suitable for short-range applications and therefore fall outside the scope of FlyZero. To support this conclusion, the authors consider a typical empty operating mass of 55% and a fuel mass fraction of 20%, and believe that a battery-powered aircraft with this energy fraction would only be able to operate with a very short range of less than 250 nautical miles (nm). The authors also conclude that the range of a battery-powered aircraft would only increase to 450 nm at the expense of reducing the payload mass fraction to zero (thus ruling out the use of the aircraft to carry passengers).

[0010] In "The challenge for battery-powered aircraft," published by Bits & Chips, 2022, H. Werij and M. Wagemaker conclude that purely battery-powered aircraft will play a very small role in reducing the climate impact of flight. The authors support this conclusion by arguing that in the expected design space for battery-powered aircraft, the empty operational mass and battery mass are approximately 60% and 25% of the maximum takeoff mass, respectively, resulting in an effective range of only 250 km.

[0011] At least two themes emerge from the existing literature. First, the ranges of existing electric aircraft and existing electric aircraft designs are so short that they cannot have a significant impact on emissions across the air transportation sector. The well-known Breguet range equation indicates that range is determined by the energy fraction (the ratio of fuel mass divided by maximum takeoff mass). However, many authors argue that this energy fraction cannot exceed 25–30% with rechargeable batteries, since current well-designed short-range fossil-fuel aircraft have similar fuel mass fractions. Second, rechargeable batteries suitable for such existing electric aircraft designs will likely not exist in the foreseeable future. According to existing electric aircraft design principles, battery energy density is simply too low.

[0012] Therefore, the common opinion in the aerospace engineering literature is that battery electric aircraft will play a very small role in achieving net zero, as the ranges obtainable using existing electric aircraft design principles are too short for electric aircraft travel and insufficient to significantly reduce emissions in the global aviation sector.

[0013] In their paper, "Technical and Environmental Assessment of an All-Electric 180-Passenger Commercial Aircraft," Progress in Aerospace Sciences, 105 (2019), 1-30, Gnadt et al. review all-electric conceptual, experimental, and commercial aircraft that have been studied over the past several decades, with a particular focus on light aircraft, along with advances in battery technology. Furthermore, they develop all-electric aircraft designs starting from an existing conventionally powered aircraft (the Airbus A320neo). The performance of these all-electric aircraft is compared with conventionally powered advanced aircraft optimized for the same short design range. The design range limits are determined for different specific energy values ​​of the aircraft. A future projection of a four-fold increase in battery pack specific energy (from a current value of 200 Wh / kg to 800 Wh / kg) would only enable flights of 500 nm. A shorter design range is found to improve energy and environmental performance.

[0014] More specifically, the Gnat et al. paper is concerned with imagining the medium- to long-term future of electric aircraft design. However, the all-electric aircraft described therein (the AEA-800 aircraft) is essentially a futuristic ideal rather than a viable aircraft design. For example, the design incorporates fictitious technology that the authors estimate will only be available in 2050. Furthermore, while the design is advanced as being suitable for commercial flight, it does not meet the critical runway and reserve energy specifications required for commercial flight. In other words, the teachings in this paper will not enable a skilled designer to develop a viable all-electric aircraft (AEA).

[0015] Furthermore, Gnat et al. is a study to assess the potential impact of battery-electric technology on the climate impact of the aviation sector. The purpose of this paper is not to provide a clean-sheet design for a new, commercially viable aircraft engine (AEA), or to provide experienced designers with a "recipe" for arriving at an effective design for a new AEA, but rather to assess the environmental impact of an A320 configured to incorporate electric propulsion (see, for example, the introduction to Section 5 of this paper). The analysis is performed for a number of different engine and battery technology scenarios. The A320 was chosen as the design basis to isolate and observe the benefits of electric propulsion applied to a modern, widely used commercial aircraft configuration.

[0016] While the above papers are very useful, in order to assess the potential impact of battery-electric aviation, several future technology assumptions are made, which is a common and valuable approach in this type of research.

[0017] However, in contrast to Gnat et al., the problem at the focus of this disclosure is how to obtain maximum range for a given technology scenario, including current technology. Gnat et al. highlights three assumptions that result in an aircraft design that is physically impossible to achieve with current and near-future technology: a. Cryogenically cooled superconducting electric motors with a power-to-weight ratio 5-10 times higher than current electric motors, which, according to the authors' literature, will only be available in 2050 (Chapter 3.2.1); b. Future aerodynamic improvements, e.g., 25% more lift during landing (Section 5.4), and c. Future improvements in structural materials, reducing structural weight by an unspecified amount.

[0018] Furthermore, the AEA-800 aircraft proposed in Gnat et al.'s paper does not meet its current specifications. The authors acknowledge that the landing runway length (3,260 meters) poses a major problem. This would make the aircraft unsuitable for many destinations, and the authors suggest, among other things, that a system such as a military aircraft carrier could be a solution. However, this does not meet the technical requirements for commercial flight. Furthermore, the aircraft's reserve energy only allows for a 30-minute flight time. Regulations stipulate that the reserve energy must not only accommodate a 30-minute "standby" reserve, but also a flight from the destination to an alternate airport, for example, to deal with bad weather at the destination. While the actual reserve reserve varies from flight to flight, the new aircraft's specifications are expected to accommodate a reserve reserve for at least 150 km of flight, which is not the case in this paper.

[0019] Furthermore, if a skilled designer were to use Gnat et al.'s AEA-800 aircraft as a starting point for a battery-electric aircraft, at least several "modifications" would need to be made to the aircraft design in order to put the aircraft design into practice, including the following: The wing size should be increased from approximately 125 m², which is required to ensure a normal landing, to approximately 180-200 m². However, this is not possible for the AEA-800 aircraft while adhering to the span constraints imposed by this paper. The non-existent HTS engine described in this paper should be replaced with a current state-of-the-art electric motor. However, this would increase the weight of the electric motor by at least five times. And, The hypothetical future materials described in this paper should replace current state-of-the-art materials, but at the expense of increased structural weight.

[0020] The effects of the above three changes can be estimated using both the data in Gnat et al.'s paper and the estimation methods presented in handbooks (e.g., Raymer). Optimistic estimates suggest that the empty weight (weight of all structures, engines, propulsion, and all systems) would increase by 12-15 tons, the battery weight would decrease to 33-36 tons, and the aerodynamic efficiency (L / D) would decrease from 18.6 to <17. This would dramatically reduce the range and leave several issues unresolved, including too low reserve energy.

[0021] In conclusion, Gnat et al.'s paper describes what happens when starting from a known design of a fossil-fuel aircraft and attempting to replace the fossil-fuel energy system with a new battery system. This approach alone will generally not lead to an AEA suitable for commercial flight. This is essentially because Gnat et al. teach away from the problem of how to design an all-electric aircraft to increase flight distance; quite the contrary, the paper assumes that an AEA can only fly relatively short distances efficiently, even taking into account future improvements in battery technology.

[0022] The integration of rechargeable batteries in electric aircraft has not been well studied. Rechargeable batteries can be recharged in situ on the aircraft, but must be replaced after reaching a maximum number of charging cycles. Therefore, the design must allow access for periodic replacement of the rechargeable batteries. However, allowing access can come at the expense of other design aspects.

[0023] It is well known that aircraft are required by aviation authorities to carry spare fuel, for example to accommodate diversions to alternate airports, and reserves for loiter flights and contingencies. Although rarely used, such spare fuel adds extra mass to the aircraft that must be carried with each flight. Summary of the Invention

[0024] The present disclosure seeks to address these and other problems encountered in the prior art by providing an improved aircraft design.

[0025] a fuselage and at least one wing arranged in a wing-body design; A plurality of thrusters; a primary energy source configured to power the thrusters, the primary energy source comprising (or optionally, and more specifically consisting of) a rechargeable battery; The aircraft has a maximum take-off mass (MTOM) of at least 8,618 kg and an electric range factor (ERF) of at least 6.

[0026] Optionally, the aircraft has an ERF of at least 6.5, at least 7, at least 7.5, at least 8, at least 8.5, or at least 9.

[0027] Optionally, the aircraft has an MTOM of at least 20,000 kg, at least 40,000 kg, at least 60,000 kg, at least 80,000 kg, or at least 100,000 kg.

[0028] Optionally, the aircraft has a maximum payload mass "PLM" of at least 3,000 kg, at least 6,000 kg, at least 9,000 kg, at least 12,000 kg, at least 15,000 kg.

[0029] Optionally, the aircraft has a maximum payload mass "PLM" that is at least 5%, at least 10%, at least 15%, at least 20%, or at least 25% of MTOM.

[0030] Optionally, the aircraft has at least 50%, at least 60%, at least 70%, at least 80%, at least 90%, or at least 95% of the total mass of rechargeable batteries associated with the wings.

[0031] Optionally, the aircraft has a lift to drag ratio "L / D" of at least 15, at least 17, or at least 19.

[0032] Optionally, the aircraft has a lift to drag ratio "L / D" between 15 and 30, between 17 and 28, or between 19 and 26.

[0033] Optionally, the aircraft further comprises a secondary energy source configured to power the propulsors, the secondary energy source having a higher energy density than the primary energy source.

[0034] Optionally, the secondary energy source comprises a liquid fuel, a non-fossil liquid fuel, a non-rechargeable battery, a metal-air battery, or an aluminum-air battery.

[0035] Optionally, the aircraft has a wing volume to fuselage volume ratio "WV / FV" of at least 0.24, at least 0.27, at least 0.30, or at least 0.33.

[0036] Optionally, the aircraft has a wing wetted area to fuselage wetted area ratio of at least 0.50, preferably at least 0.55, preferably at least 0.60, preferably at least 0.65, preferably at least 0.70, preferably at least 0.75, preferably at least 0.80, preferably at least 0.85, preferably at least 0.90, preferably at least 0.95, preferably at least 1.00, preferably at least 1.10, preferably at least 1.20, or preferably at least 1.30.

[0037] Optionally, the aircraft has a wing loading less than WL2 or in the range WL1 to WL2, where WL2 is 850 kg / m2, or preferably 800 kg / m2, or preferably 750 kg / m2, or preferably 700 kg / m2, or preferably 650 kg / m2, or preferably 600 kg / m2, or preferably 550 kg / m2, and WL1 is 100 kg / m2, or preferably 150 kg / m2, or preferably 200 kg / m2, or preferably 250 kg / m2, or preferably 300 kg / m2, or preferably 350 kg / m2, or preferably 400 kg / m2.

[0038] Optionally, the wings are attached to the underside of the fuselage and / or the wings have foldable tips.

[0039] Also, a wing assembly for a wing-and-body design aircraft, comprising: a wing structure; and a primary energy source configured to power a plurality of propulsors, the primary energy source comprising a rechargeable battery. The wing structure has a mass of at least 600 kg, and the ratio of the mass of the rechargeable battery to the mass of the wing structure is at least 4.

[0040] Optionally, the wing structure has a mass of at least 1000 kg, at least 2000 kg, at least 3000 kg, or at least 4000 kg.

[0041] Optionally, the ratio of the mass of the rechargeable battery to the mass of the wing structure is at least 4.5, at least 5, at least 5.5, or at least 6.5.

[0042] Optionally, the wing structure is configured to be attached to the underside of a fuselage, and / or The wing structure includes a foldable tip.

[0043] Optionally, the wing assembly is for a wing-and-body design aircraft having a maximum take-off mass (MTOM) of at least 8,618 kg and an electric range factor (ERF) of at least 6.

[0044] Any of the aforementioned wing assemblies may be included on any of the aforementioned aircraft.

[0045] Optionally, the plurality of thrusters includes 6, 8, 10, 12 or 14 thrusters.

[0046] Optionally, the rechargeable battery is configured to be recharged in situ.

[0047] Also provided is an aircraft comprising a fuselage and at least one wing arranged in a wing-and-body design, a plurality of propulsors, and a primary energy source configured to power the propulsors, the primary energy source comprising a rechargeable battery. The aircraft includes a secondary energy source configured to power the propulsors, the secondary energy source comprising a non-rechargeable battery having an energy density greater than the energy density of the primary energy source.

[0048] Optionally, the secondary energy source comprises one of a metal-air battery or an aluminum-air battery.

[0049] Optionally, the aircraft has a maximum take-off mass "MTOM" of at least 8,618 kg and an electric range factor "ERF" of at least 6.

[0050] Optionally, the secondary energy source is configured to operate only as a backup energy source.

[0051] Optionally, the secondary energy source is for accommodating diversions to alternate airports and necessary loiter flights and contingency backups.

[0052] Optionally, the secondary energy source has an effective energy density of greater than 400 Wh / kg, greater than 500 Wh / kg, greater than 600 Wh / kg, greater than 700 Wh / kg, or greater than 750 Wh / kg.

[0053] Also, a wing structure with load-bearing hatches; a rechargeable battery disposed within the wing structure; The load-bearing hatch is positioned to open and close an opening in the wing structure, the opening being positioned to allow access to the rechargeable battery.

[0054] Optionally, the load-bearing hatch is configured to support at least a portion of a load applied to the wing structure and / or to transfer aerodynamic, inertial or gravitational loads applied to the wing structure.

[0055] Optionally, at least a portion of the rechargeable battery is coupled to the load-bearing hatch.

[0056] Specific embodiments will now be described, by way of example only, with reference to the drawings, in which: [Brief explanation of the drawings]

[0057] [Figure 1a] 1 is a graph illustrating empty operational mass fraction versus energy mass fraction for a rechargeable battery electric aircraft, according to an embodiment. [Figure 1b] 1 is a graph illustrating empty operational mass versus energy mass fraction for a rechargeable battery electric aircraft, according to an embodiment. [Figure 1c] 1 is a graph illustrating energy consumption versus energy mass fraction for a rechargeable battery electric aircraft, according to an embodiment. [Figure 2a] 1 is a schematic graph illustrating a comparison of design spaces for a turbopropeller fossil fuel aircraft, a long-range fossil fuel aircraft, and an electric aircraft, according to an embodiment. [Figure 2b]2b shows a copy of the schematic graph shown in FIG. 2a and an adapted version of the schematic graph for an electric aircraft employing further design principles according to the present disclosure. [Figure 3] 1 is a graph illustrating maximum take-off mass "MTOM" versus battery mass fraction for a rechargeable battery aircraft, according to an embodiment. [Figure 4] FIG. 1 illustrates a schematic comparison of wingspan of an electric airplane compared to a similar turboprop airplane, according to one embodiment. [Figure 5] 1 is a graph illustrating aircraft range achieved with various parametric designs, according to an embodiment. [Figure 6] 10 is a graph illustrating aircraft MTOM obtained for various parametric designs, according to an embodiment. [Figure 7] 1 is a graph illustrating aircraft energy consumption per passenger kilometer obtained for different parametric designs (passenger numbers and ERFs), according to an embodiment. [Figure 8] FIG. 1 is a simplified plan view of a family of electric aircraft with various payloads and electric range factors. [Figure 9] FIG. 1 shows the ISO range curve in a graph of battery energy density versus ERF. DETAILED DESCRIPTION OF THE INVENTION

[0058] Overview Generally, but not by way of limitation, the present disclosure relates to defining a design space for electric aircraft that goes against conventional thinking and design trends in the field in order to produce electric aircraft with longer ranges than previously conceived using rechargeable batteries. This application also relates to wings for such electric aircraft, load-bearing hatches for aircraft wings, and non-rechargeable backup energy sources for rechargeable battery electric aircraft. "Wing" means a wing installed on an aircraft for normal use. "Wing assembly" may refer to the wing or the structure forming the wing itself, optionally together with any items, fixtures, or components associated with the wing (i.e., any items, fixtures, or components attached to, located within, or belonging to the wing). As is known, wings and wing assemblies may be initially manufactured independently and then attached to the aircraft (fuselage) at a later stage. "Wing" or "wing assembly" may identify the corresponding wing, structure, or assembly not only during or after manufacturing, but also after installation on the aircraft.

[0059] Some of the design principles in this disclosure seek to maximize useful battery mass fraction, and therefore range, by minimizing empty operational mass fraction for a given payload mass requirement, which is achieved by utilizing the design principles, choices, and features described herein.

[0060] Conventional thinking suggests that electric aircraft design will be limited to very short-range aircraft that may not be suitable for the majority of commercial flight distances. While the range of battery-powered aircraft is indeed shorter than that of fuel-based aircraft, the design principles presented herein demonstrate that the range is actually higher than literature and conventional thinking in electric aircraft design and therefore can play an important role in achieving net zero.

[0061] The solution to increasing the range of rechargeable battery aircraft is to increase what we have coined the Electric Range Factor "ERF" (lift to drag ratio times battery mass "BM" expressed as a fraction of maximum takeoff weight "MTOM") while operating in the CS-25 design space (i.e., so that the aircraft has an MTOM of at least 8618 kg).

[0062] The inventors have recognized that increasing MTOM, combined with shifting the design space to allow for increased battery mass fraction compared to existing electric aircraft designs, is a particularly effective means for increasing the electric range coefficient. The inventors have also devised several means by which the empty operational mass fraction can be reduced, thereby enabling further increases in battery mass fraction and therefore extending range. Additionally or alternatively, reducing the empty operational mass fraction allows for a larger payload mass fraction and therefore reduces energy consumption per passenger kilometer (pax-km).

[0063] The inventors have also devised design features that support these general concepts. For example, by associating battery mass primarily with the wing (e.g., by locating rechargeable batteries in, on, or under the wing), the bending moment at the wing root can be reduced. This can further reduce the empty operational mass "EOM" fraction (EOM / MTOM) because the relative mass of the aircraft structure can be reduced due to reduced stresses in flight where the wing intersects the fuselage. This is just one example of a design improvement that can enable a larger payload mass "PLM" fraction (PLM / MTOM) and / or an even larger battery mass fraction (BM / MTOM).

[0064] Embodiments employing these design principles, options, and / or features include aircraft with relatively larger wing area (and volume) and relatively smaller fuselage area (and volume) compared to a typical CS-23 or short-range CS-25 aircraft, which naturally translates into a higher lift-to-drag ratio than a typical CS-23 or CS-25 aircraft, further expanding the range for a given battery mass fraction.

[0065] Other design features or principles described in this disclosure can (further) increase the lift-to-drag ratio of an aircraft and / or enable a reduction in the empty operational mass fraction, thus enabling a larger battery mass and / or payload mass fraction, further increasing the range of an electric aircraft. For example, increasing the number of propellers compared to a conventional aircraft, such as having four or more propellers on a wing, helps increase airflow over the wing and therefore increases lift. In another example, the mass of the aircraft can be reduced by employing a low-wing structure in which the wings are connected to the underside of the fuselage. In yet another example, the wingspan and / or the size (e.g., planar area or volume) relationship between the wings and the fuselage during flight can be adjusted to increase the lift-to-drag ratio. For particularly large wingspan designs, the wings can include foldable tips to accommodate airport requirements.

[0066] Another aspect relates to a secondary energy source for an electric aircraft having a rechargeable battery as its primary energy source. The secondary energy source has a higher energy density, but due to its intended use as a reserve, it may be used very infrequently, if at all, in the normal operation of the aircraft. Thus, the aircraft can utilize a secondary energy source suited to these requirements to most efficiently utilize the mass fraction reserved for the aircraft's energy source, while allowing the aircraft to remain rechargeable by the primary energy source.

[0067] Yet another aspect relates to a load-bearing hatch in a wing structure for accessing an internal rechargeable battery. While rechargeable batteries can recharge the wing structure in situ, their useful life may be only a fraction of the wing structure's useful life, and therefore battery replacement is required from time to time. Replacement requires access into the wing structure through an opening. Such an opening potentially weakens the wing structure or reduces its ability to support loads applied to it, for example, during flight. Rather than increasing the wing structure mass by reinforcing the area around the opening to transfer applied loads around the opening, the hatch is instead configured to support the load. Thus, the wing structure mass does not need to be increased to accommodate an opening for accessing the rechargeable battery.

[0068] Detailed Description Abbreviation For the sake of brevity, the following abbreviations are used in this disclosure: BM = Battery mass DEP = Distributed Electric Propulsion (i.e., a total of four or more thrusters associated with one or more wings) EM = Energy Mass (= BM for battery electric aircraft) EOM = empty operational mass ERF = Electric Range Factor MTOM = Maximum Take-Off Mass PLM = Maximum Payload Mass PtL = Power to Liquid Fuel RPK = Revenue Passenger Kilometers SAF = Sustainable Aviation Fuel eSAF = Sustainable aviation fuel produced from electricity TLAR = Top Level Design Requirements

[0069] Symbols and Parameters The following symbols and parameters are also used: A = aspect ratio C D = drag coefficient, D / (q ∞ S)=C D0 +C Di C D0 = Zero-lift drag coefficient C Di = induced drag coefficient C L = lift coefficient, L / (q ∞ S) C Lmax = Maximum lift coefficient D=Drag force [N] D fus = fuselage diameter [m] e = Oswald coefficient e bat = Battery energy density [J / kg] g=gravitational constant: 9.8m / s 2 L / D=lift-drag ratio l fus = fuselage length [m] q ∞ = free flow pressure, 0.5ρ ∞ V 2 [Pa] R = Range [m] or [km] R fus = radius of fuselage [m] S = wing area [m 2 ] S wing / S fus = wing area to fuselage area ratio, 2S / (πl fus D fus ) V=velocity [m / s]

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[0070] definition The following definitions are useful in understanding this disclosure.

[0071] Breguet Range: Cruising flight range calculated based on the adapted Breguet range formula (see formula section).

[0072] Useful range: Maximum practical flight distance. That is, the maximum distance a given payload can be transported by an aircraft, from takeoff to landing, in normal, windless weather conditions, while respecting necessary requirements regarding reserves. This distance can be calculated by taking the Breguet range and making some adjustments for ao takeoff, climb, etc. (See calculation assumptions).

[0073] Total range: Useful range plus the effective range that an aircraft must be able to fly and respond to reserves. "Reserves" includes three major components: contingencies, diversions, and reuter flights.

[0074] Lift-to-drag ratio: The ratio between lift, defined as the component of the aerodynamic force acting on the aircraft (excluding the propellers) perpendicular to the direction of the oncoming flow, and drag, defined as the component parallel to the direction of the oncoming flow.

[0075] For existing aircraft (i.e., in the real world), this parameter can be measured, but is numerically estimated during the conceptual design process. Regarding the values ​​and limits established in this disclosure: For the purposes of measured (or real-world) aircraft, lift-to-drag ratios may be measured in steady-state flight, cruising altitude and cruising speed in ISA (International Standard Atmosphere) conditions unless otherwise stated. For the purposes of the conceptual design process, e.g., lift-drag parabolic curves

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[0076] [Table 1]

[0077] Additionally, zero-lift drag due to fuselage upsweep (Daniel Raymer's textbook, Section 12, referenced herein below) and the thermal management system heat exchangers (assuming a drag of 0.1668 N / kW of heat rejection) are included. Drag due to external stores, windmilling engines, and transonic / supersonic parasitic drag are not considered applicable and are therefore ignored.

[0078] Battery Mass (BM): The mass of a rechargeable battery pack used for propulsion and non-propulsion purposes, including energy cells, battery management systems, and elements that contribute to the structural integrity of the cells (e.g., frames or other mechanical elements), thermal management (e.g., cooling plates or channels), replaceability, or safety and reliability. In other words, a rechargeable battery pack consists of components that are removed and replaced when the cells reach the end of their life on the aircraft, and any components that remain on the aircraft during the cell replacement process but that contribute functionally (i.e., structurally, thermally, etc.) to the integrity and operability of the cells on board the aircraft.

[0079] Battery Energy Density: The sum of the end-of-life useful energy capacity of the cells in a rechargeable battery pack divided by the battery mass as defined above.

[0080] "Useful" energy refers to the maximum amount of energy that can be safely extracted from a cell during a normal discharge cycle while still meeting the specified cycle life, i.e., taking into account the maximum depth of discharge.

[0081] "End of life" refers to the cell energy capacity before it is replaced on board the aircraft, i.e., taking into account cell degradation. Note that the cells may still have secondary uses after their life on board the aircraft.

[0082] Energy capacity must be quantified at a discharge rate representative of normal operation (e.g., cruise C-rate).

[0083] For example, when new, for cells delivering 100Wh of energy per kg of cell at a discharge rate of 1C, with a 25% mass surplus for packaging, a maximum depth of discharge of 90%, and a selected end-of-life capacity of 80%, the battery energy density is (100 0.8 0.9) Wh / (1 + 0.25) kg = 57.6 Wh / kg.

[0084] Electric range factor: A dimensionless parameter defined as the product of the lift-to-drag ratio and the battery mass fraction, the latter being the ratio between the battery mass and the maximum takeoff mass of the aircraft.

[0085] Range extender: A set of powertrain components used to increase the overall range beyond that achievable with a rechargeable battery alone. For example, in the case of a fuel-based gas turbine solution (turbine generator), the "range extender" includes the gas turbine and its accessories (oil system, intake, exhaust, etc.), fuel, fuel system, generator, and the cables and other additional elements of the electrical system required to power the rechargeable battery or propulsor. In the case of a non-rechargeable battery, such as a metal-air or aluminum-air battery, the range extender comprises the energy cells, battery management system, and battery, including elements that contribute to the structural integrity (e.g., frame or other mechanical elements), thermal management (e.g., cooling plates or channels), replaceability, or safety and reliability of the cells, as well as any additional cables or power distribution elements required to transfer power from the non-rechargeable battery to the rechargeable battery or propulsor. Note that regardless of the type of range extender, the mass of the range extender is considered part of the EOM.

[0086] Primary energy source: Rechargeable batteries to power the thrusters.

[0087] Secondary Energy Source: A type of energy source other than a rechargeable battery used by the range extender. For example, a non-rechargeable battery or Jet A1 fuel. In the case of a non-rechargeable battery solution, the secondary energy source is the non-rechargeable battery itself. In the case of a fuel-based gas turbine solution, the range extender includes fuel as a second energy source and other components described herein.

[0088] Range extender effective energy density: The effective energy capacity of the range extender divided by the mass of the range extender.

[0089] "Useful" energy refers to the useful amount of energy delivered to the propeller shaft, taking into account both the available energy density of the secondary energy source and possible conversion or transmission losses between the secondary energy source and the propeller shaft.

[0090] For example, for a gas turbine-based range extender containing 500 kg of fuel with a specific energy of 43 MJ / kg, a range extender mass excluding fuel of 1000 kg, a gas turbine efficiency of 35%, and a gas turbine shaft to propeller shaft conversion / transmission efficiency of 90% (e.g., accounting for losses in cables, electric machines, and power electronics), the effective energy density is (500 43 0.35 0.9) MJ / (1000 + 500) kg = 4.52 MJ / kg = 1254 Wh / kg. For non-rechargeable battery range extenders, the effective energy density of the range extender is approximately equivalent to the "battery energy density" as defined herein, assuming that the mass of the additional power distribution elements required for the non-rechargeable battery is small compared to the mass of the non-rechargeable battery itself.

[0091] Propulsion: A device used to generate thrust, such as a propeller or fan.

[0092] "Wing-associated" battery: A battery installed to help reduce the bending moment at the wing root during steady level flight, such as a battery installed within the wing box (or wing volume) or mounted in a pod installed on or under the wing spar, wing skin, or wing.

[0093] "Rechargeable" battery: A battery (i.e., an energy storage device containing one or more electrochemical cells) that can re-store energy after being discharged by applying external power, optionally meaning that the battery can be recharged in situ while installed on board the aircraft.

[0094] Wing Structural Mass (WM): The mass of the basic wing structure, including ribs, spars, skin panels, etc., but excluding secondary elements such as high-lift devices, spoilers, speed brakes, or actuator mechanisms. For existing aircraft, this can be measured, but must be estimated in the conceptual design process. For the values ​​and limits established in this document, (basic) wing structural mass is defined according to "Synthesis of Subsonic Aircraft Design, Egbert Torenbeek, 1982, Delft University Press, Appendix C, section C-2."

[0095] Wingspan: The distance from one wingtip to the other when the wings are positioned for normal flight.

[0096] Wing: A wing has two (typically symmetrical) halves, i.e. a conventional (monoplane) aircraft has one main wing.

[0097] Battery Mass / Wing Mass Ratio (BM / WM): The mass of the batteries associated with the wing, employing the definitions of "battery mass" and "associated with the wing" given herein, whether they constitute a primary or secondary energy source, divided by the "wing structural mass" as defined herein.

[0098] Load-bearing hatch: A hatch that, when closed, acts as an integral part of the wing structure and transfers aerodynamic, inertial, gravity, or other loads from one point on the wing structure to another, maintaining the desired shape and structural integrity of the wing in flight.

[0099] formula The following equations and supporting explanations provide the basis for the calculations, assumptions, principles and concepts described in this disclosure.

[0100] The range of a battery-electric aircraft is defined by Equation 1, known as the adapted Breguet equation: R=

number

[0101] Increasing either or both of BM / MTOM and L / D is desirable to increase range. The other terms in Equation 1 are defined herein in the Symbols and Parameters section and are essentially fixed to currently achievable maximum values, as explained in the Description below. As explained in the Definitions section herein, the product of the terms BM / MTOM and L / D in Equation 1a is defined herein as the Electric Range Factor "ERF," given by Equation 1b.

number

[0102] Equation 2 defines the energy efficiency of an aircraft during cruise flight. Energy per 1 pax-km (MJ / paxkm) =

number

[0103] Increasing PLM / MTOM reduces the energy per pax-km and therefore improves energy efficiency per passenger.

[0104] Another formula is the so-called unified equation. MTOM=PLM+BM+EOM (Equation 3a)

[0105] This simply states that the total aircraft mass (or MTOM) is the sum of the maximum payload, batteries, and empty operating mass (EOM), where PLM is the maximum payload mass. This unified equation can also be written as the sum of the following three ratios:

number

[0106] Some of the design principles described herein aim to maximize BM / MTOM and / or PLM / MTOM by minimizing EOM / MTOM.

[0107] Calculation Assumptions The adapted Breguet range and energy efficiency equations (Equation 1 and Equation 2) incorporate several parameters that are the focus of the design principles of this disclosure. However, in the calculations presented herein, several values ​​must be assumed for these variables. These assumptions are representative of "large" commercial transport aircraft. The reasons for focusing on this aircraft category are explained below, particularly under the headings "Large Considerations" and "1960s Considerations."

[0108] η elec =90%. elec is the overall efficiency from battery discharge to energy delivered to the propeller shaft. It is a multiplication of four efficiencies (with assumed values ​​for calculation purposes): battery discharge efficiency (97%), cable efficiency (98%), inverter efficiency (98.5%), and electric motor efficiency (96%). The percentage values ​​quoted are high, but are achievable for these components.

[0109] η p =85%. pis the propulsive efficiency. It is a measure of the propeller's efficiency in converting the power received from the propeller shaft into thrust and speed during cruise flight. A hypothetical propeller efficiency of 85% is already achieved in current fossil-fuel turbopropeller aircraft. Distributed propulsion and optimized propeller design can improve this further.

[0110] e bat = 200-400Wh / kg. There are several factors that determine this parameter. - Energy density of individual cells (total electrical energy per kg of battery cells), Cell / pack ratio (the amount of additional material added to a battery cell to provide structural rigidity, fire protection, etc.), Maximum and minimum state of charge (batteries perform significantly differently in terms of voltage, maximum power, and internal resistance depending on their state of charge, and therefore practical minimum and maximum limits on the state of charge are often set to protect the battery from degradation); and Battery degradation (actual maximum energy density as a percentage of maximum energy density at 0 cycles).

[0111] Because useful battery energy densities in 10 years are unknown, the principles and calculations of the present disclosure may apply in the range of 200 to 400 Wh / kg. The lower end of this range represents current technology, while the upper end of this range represents targeted technology, as represented, for example, by NASA's SABERS project. Some literature uses much higher energy densities and still concludes that meaningful large-scale battery-electric aviation is not feasible. Therefore, demonstrating the feasibility of battery-electric aircraft with energy densities of 200 to 400 Wh / kg validates the design principles of the present disclosure. However, in embodiments, the rechargeable battery has an energy-to-mass ratio of at least 240 Wh / kg, preferably at least 360 Wh / kg, and more preferably at least 440 Wh / kg. In either case, in embodiments, the rechargeable battery has sufficient energy capacity to provide the aircraft with a range of at least 500 km, at least 800 km, or at least 1000 km.

[0112] L / D = 22. Most regional narrowbody aircraft have an L / D of 15-16, while current larger long-range aircraft approach an L / D of 19-20. However, as discussed in the following description, an L / D value of 20-25 can be assumed for a well-designed electric aircraft. For calculation purposes, an L / D of up to 22 was assumed.

[0113] The Breguet range (or Breguet cruise range) is calculated using these above assumptions. This range is the theoretical maximum range in cruise flight. It assumes that all available battery energy is used for propulsion and does not consider efficiency losses due to taxiing, takeoff, climb, etc.

[0114] The effective range is calculated by taking the Breguet range and applying the following corrections: Climb and descent. During climb and descent, operating conditions and efficiency are different from cruise flight. For example, propulsive efficiency during climb is 80% instead of 87%. The additional energy required for all non-propulsion systems (thermal management systems, environmental control systems, avionics, etc.) is estimated to be 6% of the propulsion energy required. Battery energy required for taxiing, takeoff and landing is assumed to be 27MJ per flight per 1000kg MTOM. For the diversion segment of aircraft equipped with gas turbine-based range extenders, the additional battery energy required to support the orbit and climb to diversion altitude is assumed to be 40 MJ per flight per 1000 kg MTOM.

[0115] Therefore, useful range represents the actual range that can be flown "from start to finish" in quiet atmosphere. Note: The calculated ranges for various parametric designs (e.g., Table 5) are useful ranges calculated as explained above.

[0116] Aircraft must carry sufficient fuel / energy for both the immediate flight and to respond to a diversion or other emergency. The required fuel (or energy) consists of the following: Ground motion energy: the amount required for starting and ground motion Flight Energy: The amount required for flight from takeoff to landing and taxiing to parking, calculated based on the expected route and altitude. Contingency energy: the amount required to cover unexpected additional flight fuel usage, e.g. due to headwinds, suboptimal routing, etc. Often 3-5% of flight fuel Alternative Energy: Fuel required in case of diversion to an alternate airport. This fuel corresponds to the diversion from the destination to the alternate location. The exact amount depends on the alternate airport selected. In some instances (clear weather and multiple runways) there is no requirement for substitution. Final reserve energy: (turbine engine) Amount sufficient for 30 minutes of loiter flight at 1500 feet at minimum required power, i.e., time spent above the airport.

[0117] The proposed design provides ground running and flight energy through rechargeable batteries. The ultimate reserve (loiter flight), alternative reserve, and contingency reserve are provided by range extenders. The total range can be calculated as the sum of the useful range and the range derived from the above three energy reserves.

[0118] Design principles for maximizing effective range Introduction and Design Principles The inventors recognize that conventional thinking in electric aircraft design assumes a very limited useful range for aircraft with rechargeable batteries. This disclosure proposes new design principles that go against existing trends, insights, and design thinking in electric aircraft.

[0119] This disclosure includes nine design principles that define a new design space for commercial battery-electric aircraft. These nine principles may be used in any combination.

[0120] 1. "Think big": Use aerodynamic and weight scale advantages.

[0121] 2. "Think 1960s": Use design principles of long-range aircraft from the 1960s.

[0122] 3. "Batteries attached to the wings": Batteries inside the wings to reduce weight.

[0123] 4. "Low power-to-weight ratio": Makes the most of the characteristics of electric motors.

[0124] 5. "Optimal Wing Loading": Uses DEP to allow longer takeoff and landing distances and lower cruise altitudes.

[0125] 6. Low-wing configuration: A low-wing design is used to carry the load most efficiently in a variety of scenarios, reducing the weight of the wings, fuselage and equipment.

[0126] 7. "Load-bearing hatch": The lightest possible wing structure to ensure access to the battery for battery changes.

[0127] 8. "Many different ways to handle reserves": Various solutions for required reserves.

[0128] 9. "High L / D is a free perk": The slenderness of normal wings and a small fuselage relative to a large wing give a high L / D.

[0129] The nine principles above may be used in any combination, although it will be appreciated that applying two or more of the nine principles in combination will provide increased synergistic effects.

[0130] "Consider large size" The Breguet range equation is independent of aircraft size. It applies to both the smallest radio-controlled aircraft and the Airbus A380. However, there are some scaling effects that affect the parameters of the Breguet equation.

[0131] 1) Some aircraft systems have a specific mass independent of the size of the aircraft. This means that the larger the aircraft, the lower the relative mass of these systems as a percentage of the total mass, and therefore the lower the relative contribution of these systems to the EOM. This applies, for example, to avionics, cockpit equipment, etc. Other systems (e.g., electrical systems) also do not increase proportionally with the size of the aircraft. In this regard, a move towards larger aircraft designs is preferable, as it allows for a reduction in the void mass fraction.

[0132] 2) The weight of an aircraft structure (fuselage, wing structure, tail section, landing gear) is governed by many factors. There are some factors that scale positively with aircraft size (i.e., larger aircraft result in a lower empty weight fraction): for example, the fuselage volume of a 100-seat aircraft is used more efficiently (fuselage volume per passenger) than a 20-seat aircraft. However, there are also factors that scale negatively and cause the empty weight fraction to increase with size, most notably the so-called "square cube law" (e.g., from Synthesis of Subsonic Aircraft Design, Egbert Torenbeek, 1982, Delft University Press). This law states that if the dimensions of an aircraft increase by a factor of 2 (length doubles, width doubles), the areas (wing area, fuselage wetted area) increase by a factor of 2. 2 The volume and therefore the weight of the structure increases by a factor of 2 3 An example of this is the wing weight fraction (wing mass divided by MTOM or WM / MTOM), which increases with wingspan if wing loading is held constant. Generally, the net effect of these two factors is a decrease in the airborne operational mass fraction as size increases. However, there is an approximate "optimum aircraft size," and for aircraft above this size, the positive scaling factor decreases, the negative scaling factor begins to dominate, and the airborne operational mass fraction begins to increase again with increasing MTOM. In practice, this point is difficult to determine precisely due to the influence of other design considerations, but a good approximation is that the optimal size is approximately 200,000 kg MTOM. While there is no technical reason for a maximum MTOM for an aircraft according to the present disclosure, it is not anticipated that an aircraft with an MTOM greater than 300,000 kg would be economically viable. Therefore, the nominal maximum MTOM for an electric aircraft according to the present disclosure is approximately 300,000 kg.

[0133] 3) Aerodynamic performance, expressed as lift-to-drag ratio (L / D), tends to increase when moving from smaller general aviation aircraft (up to 19 seats) to larger commercial aircraft. This is due to the fact that smaller aircraft relative to their size have larger frontal surfaces, less efficient fuselages, often non-retractable gear, and relatively large tail areas and control surfaces.

[0134] There are two sets of airworthiness certification requirements for passenger aircraft. In Europe, these are labeled CS-23 and CS-25. In the United States, similar rules exist. CS-23 applies to aircraft carrying 19 or fewer passengers and with a maximum weight of 8618 kg. CS-25 applies to larger aircraft. CS-25 requirements are more stringent than CS-23 requirements in many respects.

[0135] To achieve significant range, EOM / MTOM must be as low as practicable and L / D must be as high as practicable. The inventors have recognized that, given the above considerations regarding scale affecting empty operating mass, empty weight fraction, and L / D, the CS-25 design space is much more suitable for battery-electric aircraft.

[0136] Thus, in embodiments, the MTOM is greater than 8,618 kg. The MTOM may also be greater according to embodiments, with improved effects on range being realized at higher MTOMs. For example, the MTOM is greater than 20,000 kg, optionally greater than 40,000 kg, optionally greater than 60,000 kg, optionally greater than 80,000 kg, and optionally greater than 100,000 kg. The maximum MTOM may be 200,000 kg, 250,000 kg, or 300,000 kg. Thus, the range of MTOM for a rechargeable battery electric aircraft according to the present disclosure may be defined by any combination of these upper and lower limits, or any open range defined by a lower limit.

[0137] "Considering the 1960s" Previous studies on the feasibility of electric aircraft have used modern short-range fossil-fuel aircraft as a reference aircraft for comparison. Short-range fossil-fuel aircraft require relatively small amounts of fuel and therefore have a low energy mass fraction (EM / MTOM) of around 15%–20%. These aircraft typically have an empty operational mass fraction (EOM / MTOM) of around 60%. Subsequently, such studies assume that electric aircraft have, at best, a comparable EOM / MTOM because batteries are heavy (i.e., battery energy density is low compared to jet fuel energy density). However, this assumption is incorrect, and many authors have focused on the wrong part of the design space.

[0138] At the beginning of the jet age in the 1960s, several aircraft were developed with much lower empty operating mass fractions and much higher energy fractions than today's aircraft. The Boeing 707 and Douglas DC8-63 had EOM / MTOM fractions of approximately 45%. This was driven by the need to accommodate intercontinental distances, which, combined with the thermodynamic inefficiencies and conservative aerodynamics of first-generation jet engines, resulted in high fuel consumption. For example, "Aerodynamic Design of Transport Aircraft," E. Obert, IOS Press, 2009, shows the EOM / MTOM and EM / MTOM fractions for several aircraft developed between 1950 and 1990. Obert concludes that Figure 5 in his book "shows that for a given aircraft category, empty mass fraction is more or less constant and largely independent of aircraft size, but dependent on range." For long-range aircraft, the empty weight fraction is 45% and the fuel fraction is about 45%, leaving 10% for payload. Short-range aircraft have a fuel fraction of about 20-25%, an empty weight fraction of 50-60%, and a payload fraction of 25-30%. While Overt uses "weight," the reader will understand that "mass" (kg) and "weight" (N) can be used interchangeably when describing fractions or correlations of mass or weight because they differ by a constant factor g, a term that cancels out such fractions or correlations.

[0139] A similar relationship is observed in "Advanced Aircraft Design, Egbert Torenbeek, Whiley Press, 2013," where the author (Torenbeek) concludes, "This suggests that EOM fraction is more closely related to fuel fraction than to other characteristics." Torenbeek then proposes a formula for estimating Class 1 empty weight for narrow-body aircraft (all units in kg). EOM=1.25·PLM+0.2·MTOM+500 (Formula 4)

[0140] Combining this equation with the unified equation (Equation 3a) and specifying the PLM (e.g., 10,000 kg for a 100 pax aircraft), both EOM and EOM / MTOM can be expressed as a function of EM / MTOM. Furthermore, Equation 2 can be used to calculate energy consumption per passenger kilometer as a function of EM / MTOM. Equation 4 is provided for illustrative purposes only, and the relationship between EOM, PLM, and MTOM in an embodiment may or may not follow this relationship.

[0141] Figures 1a, 1b, and 1c show graphs of EOM / MTOM, EOM, and energy consumption as a function of EM / MTOM (which can also be read as BM / MTOM for a battery-electric aircraft) calculated using Equation 4, respectively. Figures 1a, 1b, and 1c, as well as Figure 3, are all for reference and illustrative purposes only and are not intended to limit the present disclosure. However, for clarity, some assumptions are listed below. EOM estimates are based on the formula from "Advanced Aircraft Design, Egbert Torenbeek, Whiley Press, 2013." EOM (kg) = 1.25 x PLM + 0.2 x MTOM + 500. PLM is assumed to be 10,000 kg, representing a 100-pax airplane.

[0142] Figures 1a, 1b and 1c show a clear relationship between the three parameters and the energy fraction EM / MTOM. High energy mass fraction (EM / MTOM) correlates with low empty operational mass fraction (EOM / MTOM). High energy mass fraction (EM / MTOM) correlates with high empty operational mass (EOM). High energy mass fraction (EM / MTOM) correlates with high energy consumption per pax-km.

[0143] These correlations are not intuitive and are often not well understood. Many authors implicitly assume that a lower EOM / MTOM means a lower EOM and therefore a more efficient aircraft. However, compared to today's aircraft, the long-range jet and propeller aircraft of the 1960s did not have a lower EOM and therefore a lower EOM / MTOM, but a higher MTOM as a result of their greater energetic mass.

[0144] A similar effect occurs for electric aircraft, where the EM / MTOM fraction is not high for long distances and is not high due to low thermodynamic efficiency, because the energy source itself is not energy dense compared to fossil fuels and is therefore very heavy with the energy required to achieve long distance flights.

[0145] In other words, when comparing an electric aircraft with a fossil fuel-based aircraft, both designed for a range of, say, 1000 km, the electric aircraft will inherently have a higher EM / MTOM and therefore an inherently lower EOM / MTOM.

[0146] To illustrate this effect, the schematic diagram shown in FIG. 2a conceptually illustrates how EOM / MTOM, PLM / MTOM, and EM / MTOM vary with range (and how EOM / MTOM and PLM / MTOM scale with EM / MTOM). FIG. 2a shows a schematic graph in which the vertical axis 201 represents mass as a percentage of MTOM and the horizontal axis 202 simultaneously represents aircraft range in km and EM / MTOM. The horizontal axis shows two scales of range: a first scale 203 for fossil-fuel aircraft and a second scale 204 for rechargeable-battery electric aircraft. The graph shows three areas: a first area 230 representing PLM / MTOM, a second area 240 representing BM / MTOM, and a third area 250 representing EOM / MTOM. The range values ​​shown in FIG. 2a are conceptual and represent typical values ​​that can be obtained using the Breguet cruise range equation.

[0147] As can be seen in Figure 2a, at the low end of the range (or EM / MTOM) scale on the horizontal axis is a first design space 210, and at the high end of the scale is a second design space 220. Note that the design spaces are merely represented schematically, and the scale of the markings representing the design space limits is not intended to limit this disclosure. With respect to fossil fuel-powered aircraft, the first design space 210 represents turboprop aircraft, and the second design space 220 represents long-range fossil fuel jet aircraft employing the "think big" and / or "think 1960s" design principles described herein.

[0148] As seen on the left side of Figure 2a, a first design space 210 is characterized by a higher empty maneuvering mass fraction, a higher payload mass fraction, and a lower energy mass fraction. As seen on the right side of Figure 2a, a second design space 220 is characterized by a lower empty maneuvering mass fraction, a lower payload mass fraction, and a higher energy mass fraction.

[0149] Table 2 compares the typical range, MTOM, EOM / MTOM, PLM / MTOM, and EM / MTOM of several aircraft with actual data to reflect the trends shown in Figure 2a. The ATR72-600, Q400, DC-8-63, and B707-320B are known fossil fuel aircraft, while the F9X represents an electric aircraft design with rechargeable batteries according to one embodiment.

[0150] It will be appreciated, therefore, that a fossil-fuel-based aircraft can be designed for either the first design space 210 (first range 203 on the order of 500-2,000 km, high payload mass fraction) or the second design space 220 (first range 203 on the order of 5,000-15,000 km, low payload mass fraction) of FIG. 2a, depending on the target market segment. However, because the range of a rechargeable battery-electric aircraft is significantly shorter than that of a fuel-based aircraft, significant mission capability can only be achieved if the range is as long as possible. The inventors have recognized that an electric aircraft designed according to the second design space 220 can achieve commercially significant mission capability. While the second design space 220 results in a lower payload mass fraction than a fossil-fuel-based aircraft designed for the same range, FIG. 1c shows that the energy consumption per passenger kilometer is still competitive due to substantially higher powertrain efficiency.

[0151] [Table 2]

[0152] The inventors recognize that a proper understanding of these scaling effects leads to an apparently simple yet ubiquitously overlooked conclusion: a well-designed electric aircraft is in a very different part of the design space than a conventional aircraft of similar range. Retrofitting an existing aircraft to have batteries, designing a new electric aircraft with the same mass fraction, or even designing a new electric aircraft for the same set of mission requirements as a short-range fuel-based aircraft will result in a suboptimal design.

[0153] This unique way of looking at the design problem for rechargeable battery-electric aircraft leads to three design options proposed in this disclosure. First, mission requirements (payload, range, takeoff distance, etc.) are not given but are treated as design variables during the exploration phase. Second, during the exploration phase, mass fractions corresponding to aircraft with similar EM / MTOM are used instead of aircraft with similar range. Third, EOM / MTOM is selected as a key parameter to minimize when making design choices at the aircraft level because it maximizes PLM / MTOM and EM / MTOM. The design principles and choices described in this disclosure under the headings "Battery Wing Association," "Low Power-to-Weight Ratio," "Optimal Wing Loading," "Low Wing Configuration," and "Load-Bearing Hatch" focus on reducing this EOM / MTOM. As shown in Figure 2b, these design principles "tilt the energy wedge down." Figure 2b shows a reproduction of Figure 2a for comparison, with the same graph type shown below for a rechargeable battery electric aircraft that uses one or more of the aforementioned principles to enable a reduction in EOM / MTOM.

[0154] If EOM / MTOM decreases and BM / MTOM (which is specifically describing a battery-electric aircraft) maintains the values ​​of the graph shown in Figure 2a, PLM / MTOM can be increased proportionally to the decrease in EOM / MTOM. Alternatively, both BM / MTOM and PLM / MTOM can be increased according to the desired objectives of the aircraft (e.g., balancing PLM and range requirements).

[0155] In summary, in terms of mass fraction, a well-designed electric aircraft should have more in common with long-range fossil-fuel aircraft from the 1960s than with modern short-range propeller aircraft. That is, for a rechargeable battery-electric aircraft according to the present disclosure, battery mass may be at least 30%, at least 35%, at least 40%, at least 45%, at least 50%, or at least 55% of MTOM. Empty operational mass may be 55% or less, 50% or less, 45% or less, 40% or less, or even 35% or less of MTOM. Payload mass may be at least 5%, at least 10%, at least 15%, at least 20%, or at least 25% of MTOM. Embodiments include any of the rechargeable battery-electric aircraft defined herein having combinations of the above-mentioned limits for BM / MTOM, EOM / MTOM, and PLM / MTOM (summing up to 100% according to the unified equation).

[0156] Even if one skilled in the art would understand that reducing empty operating mass or limiting it to extend range would be feasible, there are no existing charge-electric aircraft that achieve this because the battery mass fraction and lift-to-drag ratio are not sufficient to meet the electric range coefficient defined in this disclosure.

[0157] Another characteristic of 1960s long-range aircraft (or aircraft designed according to the second design space 220 of FIG. 2a) is that the fuselage is small compared to the size of the wings. In contrast, in electric aircraft designed using turbopropeller fossil fuel aircraft as a reference, the fuselage is large compared to the size of the wings. This is further explained under the heading "High L / D is a free perk."

[0158] "Associating the battery to the wing" The relationships between EOM, PLM and MTOM shown in Figures 2a and 2b apply to conventional fossil fuel "wing-body" aircraft that carry passengers and cargo in the fuselage and fuel / energy in the wings.

[0159] The location of fuel or batteries within an aircraft has a significant effect on the empty operational mass. Batteries do not lose their mass during flight, which is often considered a disadvantage, but in many cases can be used to reduce structural weight. Wing structural mass is primarily influenced by the wing bending moment at the wing root (wing / fuselage intersection). Fuselage structural mass is influenced by fuselage size and the weight supported within the fuselage.

[0160] The inventors have recognized that if the batteries are mounted in the fuselage rather than in the wings, the following effects occur. ·Without the weight of the battery to counteract the wing lift, the wing structural mass increases due to the increased bending moment at the wing root during horizontal flight. Increased fuselage structural mass due to the increased weight the fuselage must support (possibly increasing drag as the fuselage volume will most likely have to increase to support the batteries) A "snowball effect" occurs: as the weight of the wings and fuselage increases, more energy is required to fly the same distance, which requires more batteries, further increasing the mass of the wings and fuselage, etc. This snowball effect is particularly pronounced at high energy mass fractions.

[0161] Thus, a design principle of the present disclosure is to associate a battery with a wing according to the corresponding definitions presented herein. For example, the battery can be mounted within the wing volume (in the wing box) or housed above or below the wing. This contributes to a reduction in the wing root bending moment because the mass of the battery counteracts the contribution to the wing root bending moment from lift in steady level flight. That is, the wing root bending moment in a first direction caused by lift is partially or totally canceled by the wing root bending moment caused by the weight of the battery associated with the wing, and vice versa.

[0162] Distributing batteries along the wingspan contributes to maximizing these benefits. This principle is well known from 1950-1960s studies of so-called spanloaders; however, it was applied to flying wings with all payload within the wings (see, for example, Figure 9 in "NASA / USRA Advanced Design Program Aeronautics Division. Final Report 1988-89. Design of a Spanloader Cargo Aircraft"). These aircraft concepts, while never built, demonstrated that an empty operational mass fraction of approximately 30% was feasible. According to the present disclosure, it is the rechargeable batteries, rather than the payload, that are associated with the wings to reduce the bending moment at the wing root.

[0163] That is, even greater reduction in wing root bending moment during normal flight can be achieved by employing span-loading principles for rechargeable batteries. That is, rechargeable batteries associated with the wing can be distributed across the wingspan to reduce wing root bending moment during normal flight. As will be understood by those skilled in the art, unless specifically described as advantageous in this disclosure, specific examples of this principle are design practices based on aircraft design. The inventors' contribution is the adoption of these span-loading principles specifically for rechargeable batteries used to power the propulsors. Alternatively or additionally, span-loading principles may be employed for rechargeable batteries to reduce bending moment on the wing itself at locations other than the intersection of the wing and fuselage.

[0164] Figure 3 illustrates the magnitude of these effects by expressing MTOM as a function BM / MTOM. Figure 3 shows a graph with a horizontal axis 301 representing BM / MTOM as a percentage and a vertical axis 302 representing MTOM in kg. Figure 3 shows a first plot line 310 and a second plot line 320 representing calculations for an aircraft with batteries associated with the wings and fuselage, respectively. The calculations assume that the weights of the wings and fuselage can be estimated using established semi-empirical methods (e.g., from "Synthesis of Subsonic Aircraft Design," Egbert Torenbeek, 1982, Delft University Press), which is sufficient to initially show trends.

[0165] Figure 3 shows that the effect of associating batteries with the wings on the weight of a battery-electric aircraft is most pronounced at high battery mass fractions. For a given PLM (10,000 kg in this case), with a battery mass fraction of 45%, the MTOM increases by approximately 25% when the batteries are located in the fuselage rather than associated with the wings. This is due to the increased structural mass required to stiffen the aircraft to accommodate the higher wing root bending moments in steady level flight caused by placing the batteries in the fuselage. This directly translates to increased energy consumption and reduced energy efficiency per passenger kilometer (pax-km).

[0166] "Low power-to-weight ratio" Power-to-aircraft weight ratio is the ratio between maximum continuous engine shaft power and maximum takeoff weight (i.e., MTOM x g). Maximum continuous power is measured at sea level under standard atmospheric conditions. Electric motors (or electric engines) differ from turbine engines in three main ways.

[0167] 1. Electric motors do not suffer from "power derating": Where turbine engines lose power when air density decreases at higher altitudes and / or at higher than standard atmospheric temperatures, electric engines do not suffer from this.

[0168] 2. Electric motors have considerable capability to provide additional power beyond maximum continuous power. While turbine engines are often designed to limit their peak power output to ±5-10% of maximum continuous power, this limitation is more lenient for electric engines. At higher power ranges, electric engines become less efficient and must dissipate more heat. This impacts the size of the thermal management system, but provides ample opportunity for significantly higher non-continuous power levels for abnormal maneuvers, such as climb, go-around, or windshear maneuvers.

[0169] 3. Electric motors provide a nearly constant ratio of maximum continuous power to engine weight, regardless of motor size. Thus, ten small motors have approximately the same weight as two large motors. This allows for "distributed electric propulsion," i.e., many engines placed spanwise across the wing, which may be employed in aircraft according to the present disclosure.

[0170] Table 3 shows the maximum power difference between electric and turbine engines at different density altitude indices.

[0171] [Table 3]

[0172] Power-to-aircraft weight ratio is a key aircraft design parameter because it determines the aircraft's powertrain weight fraction and therefore EOM / MTOM. The minimum allowable power-to-aircraft weight ratio is determined by flight performance requirements. Designers can choose any power-to-aircraft weight ratio as long as the requirements are met (similar to wing loading, discussed in the next section). For gas turbine aircraft, the minimum allowable power-to-aircraft weight ratio is typically limited by one of three performance requirements: Take-off runway distance in different ambient environments. This is less limiting for electric motors due to their higher ratio of maximum instantaneous power to maximum continuous power and their less sensitivity to ambient conditions. One Engine Inoperable (OEI) minimum climb requirement. This is also less restrictive because multiple electric motors can be used without compromising efficiency or motor weight. This means, for example, that if one of 10 electric motors fails, the remaining nine motors only need to generate 11% additional power to maintain the same total power output. In contrast, if one of two gas turbines fails on a conventional aircraft, the other turbine must be able to generate 100% additional power to maintain the same total power output. All Engines Operation (AEO) climb point requirement. This requirement is also relaxed because electric motors do not exhibit a "loss of power" with altitude.

[0173] For these reasons, power-to-aircraft weight ratios of less than 50% are readily achievable for battery-electric aircraft compared to turboprop aircraft. That is, aircraft according to the present disclosure may have power-to-weight ratios of 0.2 kW / kg or less, 0.18 kW / kg or less, 0.16 kW / kg or less, or even 0.14 kW / kg or less, where weight is MTOM and power is maximum continuous power during flight.

[0174] "Optimal wing loading" The inventors have recognized that further improvements in rechargeable battery electric aircraft may be achieved by associating more than a certain percentage of rechargeable batteries with the wings. For example, an aircraft according to the present disclosure may have at least 50%, at least 60%, at least 70%, at least 80%, at least 90%, or at least 95% of its total mass in rechargeable batteries associated with the wings. This principle may be employed with or without the spanloading principle described herein.

[0175] Wing loading is the ratio of maximum takeoff mass to wing area. For civil transport aircraft, the maximum allowable wing loading (i.e., smallest possible wing size) is determined by a combination of two parameters for approach / landing conditions: the maximum lift coefficient (CLmax) and the approach speed associated with the landing distance. "Aircraft Design: a conceptual approach, 6 th Typical wing loading values ​​given in the "Aircraft Wing Loading Standards for Turbopropellers, Vol. 1, No. 1, 2018, Vol. 1, No. 1, 2018" are approximately 200-400 kg / m² for turboprop aircraft, and up to approximately 600-700 kg / m² for large jet aircraft.

[0176] It is commonly assumed that wing loading should be as large as possible, since this means that the wing size is as small as possible. For a given wing aspect ratio ("slender"), a smaller wingspan is lighter. For a given wingspan, a smaller wing is more aerodynamically efficient. For this reason, large aircraft often use complex high-lift devices to increase CLmax and thereby the maximum wing loading. Electric aircraft can further increase this CLmax by using distributed electric propellers (or DEPs as defined hereinabove) to "blow" additional air over the wings, increasing the available lift.

[0177] However, the inventors have recognized that for electric aircraft designed for commercial passenger transport, very high wing loadings are not beneficial. This is for two reasons. First, high wing loadings require more power from the engines to take off (i.e., increase the required power-to-aircraft weight ratio). For gas turbine aircraft, the benefits of high wing loadings outweigh the drawbacks of a high power-to-aircraft weight ratio, but for electric aircraft with heavy powertrains, the opposite is true: it is advantageous to have small motors, even if this means having large wings. Second, large wings provide more volume for accommodating batteries, further enabling the application of the "wing-to-battery" approach described herein.

[0178] Understanding this design principle leads to the following design choices: A relatively small wing loading is selected for an electric aircraft compared to a conventional jet aircraft operating from the same airport with a similar MTOM; a significantly larger wing loading is selected for a turboprop aircraft with a significantly smaller takeoff and landing distance requirement; this results in a wing loading value less than W2 or within the range of WL1 to WL2, where WL2 is 850 kg / m2, preferably 800 kg / m2, or preferably 750 kg / m2, or preferably 700 kg / m2, or preferably 650 kg / m2, or preferably 600 kg / m2, or preferably 550 kg / m2, and WL1 is 100 kg / m2, or preferably 150 kg / m2, or preferably 200 kg / m2, or preferably 250 kg / m2, or preferably 300 kg / m2, or preferably 350 kg / m2, or preferably 400 kg / m2, or preferably 450 kg / m2. As a result, less complex high-lift devices are needed.

[0179] "Low wing configuration" It will be understood that the aircraft described herein employ a fuselage and wings of a "wing-body" design, where the wings can be joined to the fuselage in three ways: The wings can be attached to the underside of the fuselage, known as a low-wing configuration. The wings can be attached above the fuselage, known as a high-wing configuration. The wing and fuselage can be joined through the center of the fuselage, known as a mid-wing configuration.

[0180] For passenger aircraft, the mid-wing configuration is very challenging because the structural weight increases if the wing spars are not continuous at the joint. A continuous spar would also be challenging because it would have to cross the passenger cabin partway along the fuselage. Therefore, this configuration is not used on large (>19 pax) passenger aircraft.

[0181] For conventional fossil-fuel aircraft, high-wing or low-wing configurations each have their own advantages and disadvantages, and therefore both are used. Low-wing designs generally offer a weight advantage over high-wing designs. This weight advantage arises from the fact that the landing gear is relatively short in length and can be housed in the wings. High-wing designs require either large landing gear (e.g., Fokker F27 and Fokker 50) or landing gear joined to the fuselage (e.g., ATR 72), and a robust (and heavy) structure between the wing, fuselage, and landing gear to support all the loads, especially during taxiing and landing.

[0182] Nevertheless, high-wing designs are often used for propeller aircraft. In this case, the high-wing design allows for a larger propeller diameter while maintaining the required ground clearance. Therefore, for turboprop aircraft up to 90 pax, the high-wing configuration is the dominant design.

[0183] Existing research and designs for electric or hybrid-electric aircraft typically take existing turboprop aircraft as a reference aircraft and therefore often propose high-wing configurations, however, as will be explained, this approach does not lend itself to producing rechargeable battery-electric aircraft with longer ranges.

[0184] The inventors have recognized the importance of the following four key differences between conventional fossil fuel aircraft and battery electric aircraft: For electric aircraft, the landing weight is equal to the takeoff weight, which means that the structural loads on landing are higher and heavier landing gear is often required. For electric aircraft, the energy mass fraction is extremely high, which means that "wing-associated" batteries lead to increased structural loads during landing. For electric aircraft designed in accordance with the present disclosure (e.g., employing "think big" and "think 1960s" design principles), the fuselage is small relative to the size of the wings, meaning that landing gear cannot be easily installed on the fuselage. In contrast, for electric aircraft designed using turboprop fossil fuel aircraft as a reference, the fuselage is large relative to the size of the wings. For electric aircraft, distributed propellers can achieve the same total disc area with a smaller diameter, reducing the need for ground clearance issues.

[0185] In arriving at the low wing design, the inventors also recognized the importance of the following factors: For fuselage-mounted landing gear, the fuselage frame must be heavily reinforced to transfer loads from the wings to the landing gear during landing. This increases the weight of the aircraft structure, which is especially important when batteries are located on the wings. With fuselage-mounted landing gear, the track width between the wheels is small relative to the wingspan, so a large fairing or extension mechanism is needed to position the wheels further out and avoid rollover. Therefore, in an aircraft according to the present disclosure, wing-mounted landing gear is preferred over fuselage-mounted landing gear in order to reduce the EOM / MTOM ratio. In the case of wing-mounted landing gear, the landing gear of a low-wing aircraft can be shorter and lighter than that of a high-wing aircraft due to the clearance between the underside of the wing and the ground. · Therefore, a low wing design is preferred to reduce the empty operating mass and hence EOM / MTOM of electric aircraft. Furthermore, when batteries are placed on the wings to allow for a "battery-to-wing" effect, a high wing raises the aircraft's center of gravity compared to a low wing design, requiring a larger horizontal tail moment to rotate the aircraft during takeoff. Therefore, the design principle of "low wing configuration" can be built on the design principle of "battery to wing association" and can also have a synergistic effect.

[0186] Load-bearing hatch Rechargeable batteries degrade and must be replaced at specific intervals. Depending on the aircraft's usage, a typical replacement interval of 6 to 24 months can be expected. The replacement procedure requires access to the battery, which may be located, for example, in a wing box. While battery replacement can be considered a standard procedure, it is not required regularly during line maintenance of the aircraft. Furthermore, it can be planned well in advance. Therefore, it is acceptable for the replacement procedure to take a significant amount of time (e.g., 12 to 24 hours) and it may be performed in a dedicated environment with dedicated tooling (e.g., in a hangar).

[0187] This means that load-bearing hatches, which contribute to the structural strength of the wing but require more complicated installation / removal procedures, can be used to cover the openings used to access the rechargeable batteries when they need to be replaced or maintained. Despite the potentially longer installation / removal times, load-bearing hatches are distinguished from other structural elements of the wing because they can be opened and closed multiple times as part of standard operating procedures. This opening and closing process does not require them to be destroyed, dismantled, and replaced with new hatches. It will therefore be understood that load-bearing hatches are configured to perform cycles of removal and reinstallation as part of standard operating procedures.

[0188] Current fossil-fuel aircraft also require access to the wing boxes for inspection or repair, for example, to inspect the interior of the fuel tanks. This is done periodically, in short time frames, through "access panels" that must be opened with simple tools. Therefore, these panels cannot be configured as a critical part of the wing structure; instead, all loads and stresses within the wing must be redirected around the opening. This requires a heavy aircraft structure, as additional structural reinforcement must be provided around the access panels to support the applied loads around them.

[0189] According to the present disclosure, a load-bearing hatch can instead be used to open and close the opening for battery replacement, where a removable panel is attached to the wing structure in a manner that allows the panel to support the load itself, thereby eliminating the need for reinforcing structures around the opening and allowing for a lighter overall structure.

[0190] Thus, an aircraft according to the present disclosure may include a wing assembly including a wing structure with a load-bearing hatch and a rechargeable battery disposed within the wing structure (e.g., within an interior volume), the load-bearing hatch being positioned to open and close an opening in the wing structure, the opening being positioned to allow access to the rechargeable battery.

[0191] In embodiments, the load-bearing hatch is configured to support at least a portion of a load applied to the wing structure and / or transfer aerodynamic, inertial or gravitational loads applied to the wing structure. At least a portion of the rechargeable battery may be coupled to the load-bearing hatch such that a portion of the battery is supported by the load-bearing hatch during flight and / or can be removed from the wing structure together with the load-bearing hatch during maintenance.

[0192] It should be understood that this disclosure also relates to wing assemblies having load-bearing hatches, whether or not the wing assemblies include rechargeable batteries and whether or not the wing is for an aircraft designed according to other design principles defined in this disclosure, the meaning of load-bearing hatch being defined in the Definitions section of this disclosure.

[0193] "Handling spares in various ways" The Breguet range equation also features battery energy density as an input. The "effective" energy density can be changed by using different energy sources or types. Normally, this makes no sense: the battery with the highest energy density will always give the greatest range, so mixing battery types will only reduce range.

[0194] However, the effective range of an aircraft is the Breguet range minus the range required to fly to an alternate airport, plus the so-called "final reserve range." In total, these increase the required range by ±300 to 450 kilometers (± here means "about" or "approximately"). This means that a battery-electric aircraft always carries energy over 450 kilometers, which is only used in abnormal or emergency situations. Therefore, the present disclosure proposes using this energy in the lightest possible form, since it would be very expensive and / or not rechargeable and would rarely be used.

[0195] Assuming a future battery energy density of 400 Wh / kg and an energy consumption of approximately 130 Wh / pax / km, the required spare battery weight for a 100-seater vehicle is approximately 15,000 kg. Assuming a 15 MW SAF-fuel range extender with a specific power of 5 kW / kg and 40% efficiency, the range extender weight is 3500 kg and the required spare fuel is 1000 kg. Therefore, the "effective energy density" of this SAF range extender plus fuel is ±1100 Wh / kg. This effective energy density of the SAF range extender could be equivalent to the energy density of future aluminum oxide (Al-air) non-rechargeable batteries or other metal-air batteries. These could be used as alternative range extenders to further simplify the energy system and improve reliability.

[0196] Therefore, in a rechargeable battery electric aircraft, two possibilities exist to carry the energy for the necessary reserve: the application of a conventional fossil fuel range extender used for backup and final reserve, and / or the application of aluminum-O2 (aluminum air) or other metal-air non-rechargeable batteries as a range extender replacement. In both cases, the effective energy density of the electric aircraft increases, making effective ranges of approximately 800-1000 km possible.

[0197] Thus, in accordance with the present disclosure, there is provided an aircraft comprising a fuselage and at least one wing arranged in a wing-and-body design, a plurality of propulsors, and a primary energy source configured to power the propulsors, the primary energy source comprising a rechargeable battery. The aircraft also comprises a secondary energy source configured to power the propulsors, the secondary energy source comprising a non-rechargeable battery having an energy density greater than the energy density of the primary energy source.

[0198] The secondary energy source may have an effective energy density of greater than 400 Wh / kg, greater than 500 Wh / kg, greater than 600 Wh / kg, greater than 700 Wh / kg, or greater than 750 Wh / kg. The secondary energy source is a fossil fuel, a SAF, an eSAF, or a non-rechargeable battery, such as a metal-air battery or an aluminum-air battery.

[0199] The secondary energy source may be configured to operate solely as a backup energy source, and optionally may be configured for diversion to alternate airports and to meet required loiter flight and contingency reserves.

[0200] The secondary energy source may be used only for backup purposes and not to extend useful range, although the present disclosure is not limited in this respect.

[0201] The aircraft may be designed according to one or more of the other design principles described in this disclosure, including, but not limited to, the secondary energy sources described above.

[0202] "High L / D is a free perk" Another influential parameter in the Breguet range equation is the lift-to-drag ratio: L / D. This parameter describes the amount of drag an aircraft experiences while generating the required lift. Since lift is equal to its weight in cruise flight and engine thrust is equal to drag, the ratio L / D and MTOM determine the required thrust the engine must provide in cruise flight.

[0203] However, many factors determine the ratio L / D, but for a conventional "body-and-wing" aircraft, two factors drive L / D. Aspect ratio: The ratio of the square of the wingspan to the wing area. The larger this ratio, the higher the induced drag coefficient C Di becomes smaller, so L / D becomes higher. Ratio of "wetted area" (total area "touched" by air) to wing area: The smaller the wetted area ratio, the greater the zero-lift drag coefficient C D0 becomes smaller and therefore L / D becomes higher.

[0204] Many aerodynamic improvements have been explored over the past 50 years to further increase L / D: examples include laminar flow profiles, tip thrusters, boundary layer capture, box wings, etc. However, attempts to apply such techniques to civil transport aircraft have typically yielded only moderate or no overall gains due to other impacts on the aircraft design.

[0205] Battery-electric aircraft have inherent aerodynamic advantages due to their scaling effects (e.g., those described herein with respect to engine motor size) and do not require additional complex technology. Due to the low payload-to-mass ratios achievable with other design principles described herein (e.g., "think big" and "think 1960s"), by definition, battery-electric aircraft of traditional fuselage-and-tube configuration have relatively small fuselages and large wings.

[0206] FIG. 4 compares a 100-seat battery-electric aircraft with equal wing loading (W / S) and equal wing slenderness (aspect ratio: wingspan squared divided by wing area) to a 100-seat fossil-fuel turbopropeller aircraft. The battery-electric aircraft has a much higher L / D ratio compared to the fossil-fuel aircraft due to its relatively low ratio of wetted area to wing area. This scaling effect causes the L / D ratio to exceed 20 and rise to 22-23 with increasing wingspan. Aircraft according to the present disclosure may have an in-flight wingspan (or maximum wingspan) of at least 24 meters, at least 30 meters, at least 36 meters, or at least 42 meters. To accommodate airport requirements, the wings may include foldable tips.

[0207] To put these geometric scaling effects into perspective, Table 4 compares the relative sizes of the wings and fuselage for the same aircraft shown in Table 2. The table below compares the main characteristics of the wings and fuselage in one dimension (wingspan / fuselage diameter or wingspan / fuselage length), two dimensions (wing area / fuselage area), and three dimensions (wing volume / fuselage volume). For area ratios, the ratio (2S) / (πD fus l fus ) is employed, where the numerator is approximately equal to the wetted area of ​​the wing (twice the plan view area) and the denominator is the wetted area of ​​the cylindrical fuselage (2πR fus ·l fus ) is approximately equal to

[0208] Table 4 shows that higher aspect ratios are used in short-range electric propeller aircraft than in jet aircraft of the 1960s, which results in a smaller induced drag coefficient, C Di On the other hand, the wing-area-to-body-area ratio is larger for long-range electric jet aircraft than for modern turboprop aircraft. This results in a smaller zero-lift drag coefficient, C D0 results.

[0209] Thus, an aircraft according to the present disclosure may have a wing aspect ratio greater than 8, greater than 9.5, or greater than 11, and / or a greater wing area to fuselage area ratio (i.e., more specifically, the wing-wetted area to fuselage-wetted area ratio as defined above, or an approximation thereof) of greater than 0.50, or more preferably greater than 0.55, or more preferably greater than 0.60, or more preferably greater than 0.65, or more preferably greater than 0.70, or more preferably greater than 0.75, or more preferably greater than 0.80, or more preferably greater than 0.85, or even more preferably greater than 0.90, or more preferably greater than 0.95, or even more preferably greater than 1.00. In more general terms, the mathematical product of the wing aspect ratio and the wing area to fuselage area ratio (i.e., more specifically, the wing-wetted area to fuselage-wetted area ratio as defined above, or an approximation thereof) of an aircraft according to the present disclosure may be greater than 6, greater than 8, or greater than 10.

[0210] These aspects explain how electric aircraft inherently have a lower coefficient of drag (C) than typical fuel-based aircraft. D =C D0 +C Di ), and therefore have a higher lift-to-drag ratio. Note that compared to electric aircraft, long-range jet aircraft have a smaller wingspan relative to the fuselage diameter or length, but have a comparable volume ratio. This is because jet aircraft have a lower aspect ratio (span / chord) and taper ratio (tip chord / root chord) than electric aircraft, resulting in a larger wing volume near the wing root. Thus, aircraft of the present disclosure may have a fuselage volume-to-wing volume ratio of at least 0.20, at least 0.24, at least 0.25, at least 0.27, at least 0.30, or at least 0.33. The upper limit of the wing volume-to-fuselage volume ratio is not critical (there may be no upper limit). However, the wing volume-to-fuselage volume ratio may be at most 0.40, at most 0.45, at most 0.50, at most 0.55, at most 0.60, at most 0.65, or at most 0.70.

[0211] [Table 4]

[0212] Detailed Description of Exemplary Designs and Embodiments Aircraft design often follows several steps. Starting with initial specifications, aircraft designers create an initial concept and then verify with a relatively simple set of calculations whether this concept is "in ballpark" of the Top-Level Aircraft Requirements (TLARs). The formulas and calculations that govern this conceptual verification are often labeled as "Class 1" estimates. These "Class 1" estimates are top-down (e.g., a simple estimate of the total aircraft mass based on a small number of parameters, such as payload mass and range, power required estimate, wing size estimate, etc.) and based on past experience.

[0213] As the design is further refined toward a preliminary design, designers can verify that the design meets specifications with bottom-up "Class 2" estimates. These are weight estimates for each aircraft component based on formulas derived from statistical correlations between specific parameters and the total weight of those components in existing designs. Similar techniques are available for drag estimates and related performance estimates.

[0214] Only at the end of the detailed design process can full weight and performance verification be performed. For example, the exact weight can only be determined once a wing has been designed in full detail. However, there are several intermediate verification methods available for estimating weight or drag that combine both statistical and physical models. These methods are often labeled "Class 2.5" estimates.

[0215] According to the principles described in this disclosure, battery-electric aircraft design exists in a new "design space." This, by definition, limits the use of statistical methods typically used in Class 2 design calculations. There is no prior experience with critical elements such as batteries, motors, and cooling systems. While existing Class 2 or higher estimates can be used for other aircraft components and structures, it must be noted that these represent the "fossil fuel" era, which sometimes hinders the applicability of these estimates.

[0216] Therefore, in this disclosure, the evidence provided to prove the utility of various design concepts is often based on "Class 1" estimates. In some cases, existing (fossil fuel-based) Class 1 estimates have been adjusted, and these adjustments have been extrapolated through additional calculations. The final section provides an initial set of "Class 2" level calculations to prove the applicability of the full set of design principles for maximizing the range of battery-electric aircraft.

[0217] Therefore, Class 1 methods are used for all calculations in the parametric studies of this disclosure, e.g., the relationships between EOM / MTOM and BM / MTOM, EOM and BM / MTOM, and energy efficiency (Wh / pax-km) and BM / MTOM.

[0218] The following calculations and examples use an improved "Class 2" analysis to calculate the weight, weight fraction, aerodynamic performance (L / D) and resulting range and energy efficiency of aircraft structures and systems comparable to fossil fuel aircraft, as well as some simple estimates of other systems, based on established handbook methods and assumptions about new technologies (e.g., power-to-weight ratios of electric motors, weight estimates for heat pumps, etc.). See "Aerodynamic Design of Transport Aircraft," E. Obert, 2009; "Aircraft Design: A Conceptual Approach," 6 thedition, Daniel Raymer, 2018, and for further details of the methods in these handbooks see Synthesis of Subsonic Aircraft Design, Egbert Torenbeek, 1982, Delft University Press.

[0219] Explanation of "Prior Art" According to recent consulting reports, e.g., "Investigating the Commercial Potential of Battery Electric Aviation and Mapping Autonomous Urban Air Mobility's Progress," Roland Berger Inc., 2017 and 2019, there are over 100 new startups working on electric aircraft. Most are in the eVTOL space, electric vertical take-off and landing aircraft for intra-urban transport of 4-9 passengers. There are also several general aviation efforts aimed at recreational flying and flight training (e.g., Bye Aerospace). The only EASA-certified electric aircraft currently on the market is the Pipistrel Velis Electro two-seater trainer. Known efforts on battery-electric aircraft for passenger transport in the CS-23 design space include: Eviation Alice: A nine-seat aircraft entirely battery-powered with an estimated 400km range (including reserves). Battery energy density assumptions unknown. BM / MTOM mass claimed to be >40%. Vaeridion: A nine-seat aircraft that will be fully battery-powered and have an estimated range of 550km (including reserves). Battery energy density assumptions are unknown.

[0220] Known efforts for battery-electric aircraft for passenger transport in the CS-25 design space include: Heart Aerospace ES-30: A 30-seat battery-powered aircraft with a turbine range extender to meet backup requirements. Estimated battery range is 200 km. Batteries are located inside the fuselage. BM / MTOM < 25% and MTOM 20,000-21,000 kg. Battery energy density unknown. Maeve Echelon 1: Battery-powered aircraft with 40+ seats. Estimated battery range of 500km is assumed to include spares. Batteries are located inside the fuselage. MTOM 45,000kg and PLM ±5000kg. BM unknown. Assumptions regarding battery energy density unknown.

[0221] There have also been some efforts to electrify existing aircraft designs, i.e., by retrofitting existing airframes with new battery-powered propulsion systems. However, these efforts were abandoned when it became apparent that this approach had physical limitations. A simple illustration based on an ATR72 turbopropeller aircraft illustrates these limitations:

[0222] The ATR72 can carry ±70 pax (7400 kg) and a maximum of 5000 kg of fuel. The empty weight is 13600 kg and the MTOM is 23000 kg. Note that when carrying maximum passenger weight, the aircraft cannot carry maximum fuel, as this would violate the MTOM. ·ATR72 has an L / D of ±16. When fitted with a 30% MTOM (6900 kg) battery pack, the aircraft can carry 25 passengers (2500 kg) to stay within the 23000 kg MTOM limit. This assumes an electric engine of similar weight to the current turboprop engine, as well as power mgt and thermal mgt systems. Assuming a battery energy density of 400Wh / kg (relatively high by current standards), the total range would be 540km including the reserve. This implies a practical range of ±200km, which would not be commercially viable. Energy efficiency would be at least ±210Wh / pax km.

[0223] Parametric characterization studies based on applied design principles As an example, several parametric designs are presented herein for an aircraft including a fuselage and wings arranged in a wing-and-body design, a plurality of propulsors, and a primary energy source configured to power the propulsors, the primary energy source comprising a rechargeable battery. The aircraft has a maximum takeoff mass (MTOM) of at least 8,618 kg and an electric range factor (ERF) of at least 6. The aircraft may or may not include any of the features described based on any or all of the above design principles, or any combination of two or more principles.

[0224] Three aircraft "sizes" are evaluated for a range of ERF values: 40 passengers, 80 passengers, and 120 passengers. The aircraft is designed for a runway length of approximately 2000 meters (m), a cruise altitude of 7000 m, and a cruise speed of Mach 0.6. A wing loading of MTOM / S = 517 kg / m2 is selected to meet performance requirements, and the number and size of propellers are adapted to maintain constant disk loads in various configurations. A wing aspect ratio of 12 is selected. A 150 km diversion is assumed for reserves, with a 30-minute loiter flight capability and an additional 5% contingency reserve. These reserves are accommodated by a fuel-based range extender.

[0225] Because battery energy density is largely determined by external factors unrelated to aircraft design, for this comparison a useful end-of-life packet level battery energy density of 292Wh / kg is assumed. Because aircraft range scales linearly with battery energy density, the results can be easily extrapolated to other battery technology levels.

[0226] The range, MTOM, energy consumption, empty operational mass fraction, battery mass fraction, lift-to-drag ratio, wing mass (excluding flaps and other secondary elements), and battery mass / wing mass fraction obtained for the various parametric designs are shown in Table 5 and Figures 5 to 7.

[0227] [Table 5]

[0228] Figure 5 shows how the range of an electric aircraft varies linearly with ERF, independent of passenger load, as predicted by the Breguet equation. Furthermore, Figure 6 shows how 12 designs fit within the design space of claim 1 in terms of both ERF and MTOM, with MTOM increasing with both passenger load and ERF (i.e., range). The MTOM values ​​obtained for these designs range from typical turboprop values ​​(20t-30t) to narrowbody values ​​(70t-90t) and beyond (for purposes of this disclosure, 1 ton is 1000 kg). Furthermore, Figure 7 shows how energy consumption per passenger-kilometer improves with increasing passenger load ("considering large"). The difference between 80 and 120 passenger loads is much smaller than the difference between 40 and 80 passenger loads, indicating that energy efficiency per passenger-kilometer approaches a plateau above 80 passengers. Furthermore, energy consumption is less at lower ERF values, which corresponds to a lower battery mass fraction but a higher payload mass fraction.

[0229] 8 shows a schematic graph illustrating the design space for passenger battery-electric aircraft, according to an embodiment. The graph is defined in terms of electric range factor on horizontal axis 801 and number of passengers on vertical axis 802. Plan view diagrams of a first aircraft 810, a second aircraft 820, a third aircraft 830, a fourth aircraft 840, and a fifth aircraft 850, according to an embodiment, appear across the design space. Design parameters for each of the first through fifth aircraft are shown in Table 6.

[0230] [Table 6]

[0231] The drawing shows schematically how fuselage size increases with payload (in this case passenger count) only, while wing size increases with both payload and ERF (i.e., range). The outline of the intermediate configuration (80 passengers, ERF=9) is shown in grey, superimposed on the other four configurations for comparison.

[0232] In Figure 8, where the comparison is made at constant aspect ratio and wing loading, the wingspan b is

number

[0233] In summary, these design examples demonstrate that the design principles described in this disclosure are applicable to a wide range of payload, range, and other top-level aircraft requirements while ensuring a technically feasible and energy-efficient aircraft.

[0234] As described herein and illustrated by the embodiments and examples, the proposed design principles primarily affect BM / MTOM and also increase L / D "as a free benefit." Therefore, the merit parameter for the proposed claims is the product of these two dimensionless parameters:

number

[0235] Note: R is in kilometers, e bat The unit of is kJ / kg. Figure 9 shows the two-dimensional ERF vs. e bat9 plots the Breguet cruise range, it will be appreciated that exemplary aircraft ranges are calculated based on a more accurate fit to the Breguet equation.

[0236] FIG. 9 clearly displays what ranges are feasible in the design space (ERF greater than 6) depending on battery energy density and ERF. It is already known that, all other things being equal, the higher the battery energy density, the longer the range of an electric aircraft. However, a contribution of this disclosure is the recognition that range can be significantly increased by designing the aircraft to increase the ERF. The higher the ERF, the lower the battery energy density required to achieve a particular range. By employing any of the design principles, design features, or embodiments described herein, the range, and therefore the feasibility, of electric aircraft for commercial flight can be significantly improved.

[0237] It should be understood that the above description is intended to be illustrative, and not limiting. Many other embodiments will be apparent to those skilled in the art upon reading and understanding the above description. While the present disclosure has been described with reference to certain exemplary implementations, it will be recognized that the present disclosure is not limited to the described implementations, but can be practiced with modification and alteration within the scope of the appended claims. Accordingly, the specification and drawings should be regarded in an illustrative, and not a restrictive, sense.

[0238] The following numbered embodiments are also described herein:

[0239] Embodiment 1. A fuselage and at least one wing arranged in a wing-body design; A plurality of thrusters; a primary energy source configured to power the propulsors, the primary energy source comprising a rechargeable battery; An aircraft comprising: A maximum take-off mass (MTOM) of at least 8,618 kg; an electric range factor (ERF) of at least 6; An aircraft having:

[0240] Embodiment 2. The aircraft of embodiment 1, having an ERF of at least 6.5.

[0241] Embodiment 3. The aircraft of embodiment 1 or 2, having an ERF of at least 7.

[0242] Embodiment 4. The aircraft of any preceding embodiment, having an ERF of at least 7.5.

[0243] Embodiment 5. The aircraft of any preceding embodiment, having an ERF of at least 8.

[0244] Embodiment 6. The aircraft of any preceding embodiment, having an ERF of at least 8.5.

[0245] Embodiment 7. The aircraft of any preceding embodiment, having an ERF of at least 9.

[0246] Embodiment 8. The aircraft of any preceding embodiment, having an MTOM of at least 20,000 kg.

[0247] Embodiment 9. The aircraft of any preceding embodiment, having an MTOM of at least 40,000 kg.

[0248] Embodiment 10. The aircraft of any preceding embodiment, having an MTOM of at least 60,000 kg.

[0249] Embodiment 11. The aircraft of any preceding embodiment, having an MTOM of at least 80,000 kg.

[0250] Embodiment 12. The aircraft of any preceding embodiment, having an MTOM of at least 100,000 kg.

[0251] Embodiment 13. The aircraft of any preceding embodiment, having a maximum payload mass "PLM" of at least 3,000 kg.

[0252] Embodiment 14. The aircraft of any preceding embodiment, having a maximum payload mass "PLM" of at least 6,000 kg.

[0253] Embodiment 15. The aircraft of any preceding embodiment, having a maximum payload mass "PLM" of at least 9,000 kg.

[0254] Embodiment 16. The aircraft of any preceding embodiment, having a maximum payload mass "PLM" of at least 12,000 kg.

[0255] Embodiment 17. The aircraft of any preceding embodiment, having a maximum payload mass "PLM" of at least 15,000 kg.

[0256] Embodiment 18. The aircraft of any preceding embodiment, having a maximum payload mass "PLM" of at least 5% of MTOM.

[0257] Embodiment 19. The aircraft of any preceding embodiment, having a maximum payload mass "PLM" of at least 10% of MTOM.

[0258] Embodiment 20. The aircraft of any preceding embodiment, having a maximum payload mass "PLM" of at least 15% of MTOM.

[0259] Embodiment 21. The aircraft of any preceding embodiment, having a maximum payload mass "PLM" of at least 20% of MTOM.

[0260] Embodiment 22. The aircraft of any preceding embodiment, having a maximum payload mass "PLM" of at least 25% of MTOM.

[0261] Embodiment 23. The aircraft of any preceding embodiment, having at least 50% of the total mass of rechargeable batteries associated with at least one wing.

[0262] Embodiment 24. The aircraft of any preceding embodiment, having at least 60% of the total mass of rechargeable batteries associated with at least one wing.

[0263] Embodiment 25. The aircraft of any preceding embodiment, having at least 70% of the total mass of rechargeable batteries associated with at least one wing.

[0264] Embodiment 26. The aircraft of any preceding embodiment, having at least 80% of the total mass of rechargeable batteries associated with at least one wing.

[0265] Embodiment 27. The aircraft of any preceding embodiment, having at least 90% of the total mass of rechargeable batteries associated with at least one wing.

[0266] Embodiment 28. An aircraft according to any one of embodiments 1 to 27, having at least 95% of the total mass of rechargeable batteries associated with at least one wing.

[0267] Embodiment 29. The aircraft of any preceding embodiment, having a lift-to-drag ratio "L / D" of at least 15.

[0268] Embodiment 30. The aircraft of any preceding embodiment, having a lift-to-drag ratio "L / D" of at least 17.

[0269] Embodiment 31. The aircraft of any preceding embodiment, having a lift-to-drag ratio "L / D" of at least 19.

[0270] Embodiment 32. The aircraft of any preceding embodiment, having a lift-to-drag ratio "L / D" of at least 21.

[0271] Embodiment 33. The aircraft further comprises a secondary energy source configured to power the propulsor, the secondary energy source having a higher energy density than the primary energy source; Optionally, the secondary energy source comprises a liquid fuel, a non-fossil liquid fuel, a non-rechargeable battery, a metal-air battery, or an aluminum-air battery.

[0272] Embodiment 34. The aircraft of any preceding embodiment, having a wing volume to fuselage volume ratio "WV / FV" of at least 0.24.

[0273] Embodiment 35. The aircraft of any preceding embodiment, having a wing volume to fuselage volume ratio "WV / FV" of at least 0.27.

[0274] Embodiment 36. The aircraft of any preceding embodiment, having a wing volume to fuselage volume ratio "WV / FV" of at least 0.30.

[0275] Embodiment 37. The aircraft of any preceding embodiment, having a wing volume to fuselage volume ratio "WV / FV" of at least 0.33.

[0276] Embodiment 38. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 0.50.

[0277] Embodiment 39. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 0.55.

[0278] Embodiment 40. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 0.60.

[0279] Embodiment 41. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 0.65.

[0280] Embodiment 42. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 0.70.

[0281] Embodiment 42i. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 0.75.

[0282] Embodiment 42ii. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 0.80.

[0283] Embodiment 42iii. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 0.85.

[0284] Embodiment 42iv. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 0.90.

[0285] Embodiment 42v. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 0.95.

[0286] Embodiment 42vi. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 1.00.

[0287] Embodiment 42vii. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 1.10.

[0288] Embodiment 42viii. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 1.20.

[0289] Embodiment 42ix. The aircraft of any preceding embodiment, having a wing wetted area to fuselage wetted area ratio of at least 1.30.

[0290] Embodiment 42x. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 850 kg / m2 and WL1 is 200 kg / m2.

[0291] Embodiment 42xi. An aircraft as described in any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 800 kg / m2 and WL1 is 200 kg / m2.

[0292] Embodiment 42xii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 750 kg / m2 and WL1 is 200 kg / m2.

[0293] Embodiment 42xiii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 700 kg / m2 and WL1 is 200 kg / m2.

[0294] Embodiment 42xiv. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 650 kg / m2 and WL1 is 200 kg / m2.

[0295] Embodiment 42xv. The aircraft of any preceding embodiment having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 600 kg / m2 and WL1 is 200 kg / m2.

[0296] Embodiment 42xvi. The aircraft of any preceding embodiment having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 550 kg / m2 and WL1 is 200 kg / m2.

[0297] Embodiment 42xvii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 850 kg / m2 and WL1 is 250 kg / m2.

[0298] Embodiment 42xviii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 800 kg / m2 and WL1 is 250 kg / m2.

[0299] Embodiment 42xix. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 750 kg / m2 and WL1 is 250 kg / m2.

[0300] Embodiment 42xx. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 700 kg / m2 and WL1 is 250 kg / m2.

[0301] Embodiment 42xxi. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 650 kg / m2 and WL1 is 250 kg / m2.

[0302] Embodiment 42xxii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 600 kg / m2 and WL1 is 250 kg / m2.

[0303] Embodiment 42xxiii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 550 kg / m2 and WL1 is 250 kg / m2.

[0304] Embodiment 42xxiv. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 850 kg / m2 and WL1 is 300 kg / m2.

[0305] Embodiment 42xxv. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 800 kg / m2 and WL1 is 300 kg / m2.

[0306] Embodiment 42xxvi. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 750 kg / m2 and WL1 is 300 kg / m2.

[0307] Embodiment 42xxvii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 700 kg / m2 and WL1 is 300 kg / m2.

[0308] Embodiment 42xxviii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 650 kg / m2 and WL1 is 300 kg / m2.

[0309] Embodiment 42xxix. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 600 kg / m2 and WL1 is 300 kg / m2.

[0310] Embodiment 42xxx. The aircraft of any preceding embodiment having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 550 kg / m2 and WL1 is 300 kg / m2.

[0311] Embodiment 42xxxi. The aircraft of any preceding embodiment having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 850 kg / m2 and WL1 is 350 kg / m2.

[0312] Embodiment 42xxxii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 800 kg / m2 and WL1 is 350 kg / m2.

[0313] Embodiment 42xxxiii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 750 kg / m2 and WL1 is 350 kg / m2.

[0314] Embodiment 42xxxiv. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 700 kg / m2 and WL1 is 350 kg / m2.

[0315] Embodiment 42xxxv. The aircraft of any preceding embodiment having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 650 kg / m2 and WL1 is 350 kg / m2.

[0316] Embodiment 42xxxvi. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 600 kg / m2 and WL1 is 350 kg / m2.

[0317] Embodiment 42xxxvii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 550 kg / m2 and WL1 is 350 kg / m2.

[0318] Embodiment 42xxxviii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 850 kg / m2 and WL1 is 400 kg / m2.

[0319] Embodiment 42xxxix. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 800 kg / m2 and WL1 is 400 kg / m2.

[0320] Embodiment 42x1. The aircraft of any preceding embodiment having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 750 kg / m2 and WL1 is 400 kg / m2.

[0321] Embodiment 42xli. The aircraft of any preceding embodiment having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 700 kg / m2 and WL1 is 400 kg / m2.

[0322] Embodiment 42xlii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 650 kg / m2 and WL1 is 400 kg / m2.

[0323] Embodiment 42xliii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 600 kg / m2 and WL1 is 400 kg / m2.

[0324] Embodiment 42xliv. An aircraft as described in any preceding embodiment having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, where WL2 is 550 kg / m2 and WL1 is 400 kg / m2.

[0325] Embodiment 42xlv. An aircraft as described in any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, where WL2 is 850 kg / m2 and WL1 is 100 kg / m2.

[0326] Embodiment 42xlvi. An aircraft as described in any preceding embodiment having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, where WL2 is 800 kg / m2 and WL1 is 100 kg / m2.

[0327] Embodiment 42xlvii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 750 kg / m2 and WL1 is 100 kg / m2.

[0328] Embodiment 42xlviii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 700 kg / m2 and WL1 is 100 kg / m2.

[0329] Embodiment 42xlix. An aircraft as described in any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, where WL2 is 650 kg / m2 and WL1 is 100 kg / m2.

[0330] Embodiment 421. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 600 kg / m2 and WL1 is 100 kg / m2.

[0331] Embodiment 421. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 550 kg / m2 and WL1 is 100 kg / m2.

[0332] Embodiment 42lii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 850 kg / m2 and WL1 is 150 kg / m2.

[0333] Embodiment 42liii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 800 kg / m2 and WL1 is 150 kg / m2.

[0334] Embodiment 42 liv. An aircraft as described in any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, where WL2 is 750 kg / m2 and WL1 is 150 kg / m2.

[0335] Embodiment 42. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 700 kg / m2 and WL1 is 150 kg / m2.

[0336] Embodiment 42lvi. An aircraft according to any preceding embodiment, having a wing loading less than WL2 or in the range of WL1 to WL2, where WL2 is 650 kg / m2 and WL1 is 150 kg / m2.

[0337] Embodiment 42lvii. The aircraft of any preceding embodiment, having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 600 kg / m2 and WL1 is 150 kg / m2.

[0338] Embodiment 42lviii. The aircraft of any preceding embodiment having a wing loading less than WL2 or a wing loading in the range of WL1 to WL2, wherein WL2 is 550 kg / m2 and WL1 is 150 kg / m2.

[0339] Embodiment 43. At least one wing is attached to the underside of the fuselage, and / or

[0023] The aircraft of any preceding embodiment, wherein at least one wing comprises a foldable tip.

[0340] Embodiment 44. The aircraft of any preceding embodiment, having a power-to-weight ratio of 0.2 kW / kg or less.

[0341] Embodiment 45. The aircraft of any preceding embodiment, having a power-to-weight ratio of 0.18 kW / kg or less.

[0342] Embodiment 46. The aircraft of any preceding embodiment, having a power-to-weight ratio of 0.16 kW / kg or less.

[0343] Embodiment 47. The aircraft of any preceding embodiment, having a power-to-weight ratio of 0.14 kW / kg or less.

[0344] Embodiment 48. The aircraft of any preceding embodiment, having a secondary energy source configured to power the propulsors, the secondary energy source comprising a non-rechargeable battery having an energy density greater than the energy density of the primary energy source.

[0345] Embodiment 49. The secondary energy source comprises: Metal-air battery, or Aluminum-air battery 49. An aircraft as described in embodiment 48, comprising one of:

[0346] Embodiment 50. A wing assembly for a body-and-tube design aircraft, comprising: The wing structure and a primary energy source configured to power the plurality of propulsors, the primary energy source comprising a rechargeable battery; Equipped with A wing assembly, wherein the wing structure has a mass of at least 600 kg and the ratio of the mass of the rechargeable battery to the mass of the wing structure is at least 4.

[0347] Embodiment 51. A wing assembly as described in embodiment 50, wherein the wing structure has a mass of at least 1000 kg.

[0348] Embodiment 52. A wing assembly as described in embodiment 50 or 51, wherein the wing structure has a mass of at least 2000 kg.

[0349] Embodiment 53. A wing assembly described in any of embodiments 50 to 52, wherein the wing structure has a mass of at least 3000 kg.

[0350] Embodiment 54. A wing assembly described in any of embodiments 50 to 53, wherein the wing structure has a mass of at least 4000 kg.

[0351] Embodiment 55. A wing assembly described in any of embodiments 50 to 54, wherein the ratio of the mass of the rechargeable battery to the mass of the wing structure is at least 4.5.

[0352] Embodiment 56. A wing assembly described in any of embodiments 50 to 54, wherein the ratio of the mass of the rechargeable battery to the mass of the wing structure is at least 5.

[0353] Embodiment 57. A wing assembly described in any of embodiments 50 to 54, wherein the ratio of the mass of the rechargeable battery to the mass of the wing structure is at least 5.5.

[0354] Embodiment 58. A wing assembly described in any of embodiments 50 to 54, wherein the ratio of the mass of the rechargeable battery to the mass of the wing structure is at least 6.

[0355] Embodiment 59. The wing structure is configured to be attached to the underside of the fuselage, and / or 59. A wing assembly according to any one of embodiments 50 to 58, wherein the wing structure comprises a foldable tip.

[0356] Embodiment 60. A wing assembly according to any of embodiments 50 to 59, wherein the wing structure comprises a load-bearing hatch, a rechargeable battery is disposed within the wing structure, the load-bearing hatch is disposed to open and close an opening in the wing structure, and the opening is disposed to allow access to the rechargeable battery.

[0357] Embodiment 61. A wing assembly as described in embodiment 60, wherein the load-bearing hatch is configured to support at least a portion of the load applied to the wing structure and / or to transmit aerodynamic, inertial or gravitational loads applied to the wing structure.

[0358] Embodiment 62. A wing assembly as described in embodiment 60 or 61, wherein at least a portion of the rechargeable battery is coupled to the load-bearing hatch.

[0359] Embodiment 63. A wing assembly as described in any of embodiments 50 to 62, for a wing-and-body design aircraft having a maximum takeoff mass (MTOM) of at least 8,618 kg and an electric range factor (ERF) of at least 6.

[0360] Embodiment 64. An aircraft as described in any one of embodiments 1 to 49, comprising a wing assembly as described in any one of embodiments 50 to 62.

[0361] Embodiment 65. An aircraft as described in any one of embodiments 1 to 49 or 64, or a wing assembly as described in any one of embodiments 50 to 62, wherein the plurality of propulsors includes six propulsors.

[0362] Embodiment 66. An aircraft described in any one of embodiments 1 to 49 or 64, or a wing assembly described in any one of embodiments 50 to 62, wherein the plurality of propulsors includes eight propulsors.

[0363] Embodiment 67. An aircraft described in any one of embodiments 1 to 49 or 64, or a wing assembly described in any one of embodiments 50 to 62, wherein the plurality of propulsors includes 10 propulsors.

[0364] Embodiment 68. An aircraft described in any one of embodiments 1 to 49 or 64, or a wing assembly described in any one of embodiments 50 to 62, wherein the plurality of propulsors includes 12 propulsors.

[0365] Embodiment 69. An aircraft described in any one of embodiments 1 to 49 or 64, or a wing assembly described in any one of embodiments 50 to 62, wherein the rechargeable battery is configured to be recharged in situ.

[0366] Embodiment 70. The aircraft of any one of embodiments 1 to 49, having a maximum takeoff mass (MTOM) of 20 to 150 tons.

[0367] Embodiment 71. The aircraft of any one of embodiments 1 to 49 or 70, having a wing volume to fuselage volume ratio (WV / FV) of 0.20 to 0.50.

[0368] Embodiment 72. The aircraft of embodiment 70 having an MTOM of 50 to 100 tons.

[0369] Embodiment 73. The aircraft of embodiment 71 having a WV / FV of 0.25 to 0.40.

[0370] Embodiment 74. The aircraft of any of embodiments 1-49 or 70-73, having a lift / drag ratio (L / D) of 20-30, optionally a lift / drag ratio (L / D) of 20-23.

[0371] Embodiment 75. An aircraft described in any of embodiments 1 to 49 or 70 to 74, having a battery-powered range of at least 500 km.

[0372] Embodiment 76. The aircraft of embodiment 75, having a battery-powered range of at least 800 km.

[0373] Embodiment 77. An aircraft according to embodiment 75, having a battery-powered range of at least 1000 km.

[0374] The following numbered items are also described herein:

[0375] Item 1. A fuselage and at least one wing arranged in a wing-body design; A plurality of thrusters; a primary energy source configured to power the propulsors, the primary energy source comprising a rechargeable battery; a secondary energy source configured to power the propulsors, the secondary energy source comprising a non-rechargeable battery having an energy density greater than an energy density of the primary energy source; An aircraft comprising:

[0376] Item 2. Secondary energy sources are: Metal-air battery, or Aluminum-air battery Item 1. The aircraft according to item 1,

[0377] Item 3. The aircraft of items 1, 2, or 3, wherein the secondary energy source is configured to operate only as a backup energy source, optionally for diversion to an alternate airport and to meet required loiter flight and contingency reserves.

[0378] Item 4. The aircraft according to any one of Items 1 to 4, wherein the secondary energy source has an effective energy density of greater than 400 Wh / kg.

[0379] Item 5. The aircraft according to any one of Items 1 to 4, wherein the secondary energy source has an effective energy density of greater than 500 Wh / kg.

[0380] Item 6. The aircraft according to any one of Items 1 to 4, wherein the secondary energy source has an effective energy density of greater than 600 Wh / kg.

[0381] Item 7. The aircraft according to any one of Items 1 to 4, wherein the secondary energy source has an effective energy density of greater than 700 Wh / kg.

[0382] Item 8. The aircraft according to any one of Items 1 to 4, wherein the secondary energy source has an effective energy density of greater than 750 Wh / kg.

[0383] Item 9. Maximum take-off mass (MTOM) of at least 8,618 kg; an electric range factor (ERF) of at least 6; 10. The aircraft of any preceding item, having

[0384] Item 10. The aircraft of item 9, having an ERF of at least 6.5, at least 7, at least 7.5, at least 8, at least 8.5, or at least 9.

[0385] Item 11. The aircraft according to items 9 or 10, having an MTOM of at least 20,000 kg, at least 40,000 kg, at least 60,000 kg, at least 80,000 kg, or at least 100,000 kg.

[0386] Item 12. The aircraft according to any one of Items 9 to 11, having a lift-to-drag ratio "L / D" of at least 15, at least 17, at least 19, or at least 21.

[0387] Item 13. The aircraft according to any one of Items 9 to 12, having a wing volume to fuselage volume ratio “WV / FV” of at least 0.24, at least 0.27, at least 0.30, or at least 0.33, or a wing area to fuselage area ratio of at least 0.5, at least 0.7, at least 0.9, at least 1.1, or at least 1.3.

[0388] Item 14. The aircraft further comprises a secondary energy source configured to power the thruster, the secondary energy source having a higher energy density than the primary energy source; 14. The aircraft of any of items 9 to 13, optionally wherein the secondary energy source comprises a liquid fuel, a non-fossil liquid fuel, a non-rechargeable battery, a metal-air battery, or an aluminum-air battery.

[0389] Item 15. The aircraft of any preceding item, wherein the rechargeable battery is configured to be recharged in situ.

[0390] The following numbered clauses are also set forth herein:

[0391] Article 1. Wing structure with load-bearing hatches; A rechargeable battery and 1. A wing assembly comprising: Rechargeable batteries are placed within the wing structure, Load-bearing hatches are positioned to open and close openings in the wing structure; The wing assembly, wherein the opening is configured to allow access to the rechargeable battery.

[0392] Clause 2. A wing assembly as described in clause 1, wherein the load-bearing hatch is configured to support at least a portion of the load applied to the wing structure and / or to transmit aerodynamic, inertial or gravitational loads applied to the wing structure.

[0393] Clause 3. A wing assembly as described in clause 1 or 2, wherein at least a portion of the rechargeable battery is coupled to the load-bearing hatch.

[0394] Clause 4. A wing assembly as described in any preceding clause, wherein the wing structure has a mass of at least 600 kg and the ratio of the mass of the rechargeable battery to the mass of the wing structure is at least 4.

[0395] Clause 5. A wing assembly according to any preceding clause having a mass of at least 1,000 kg, at least 2,000 kg, at least 3,000 kg, or at least 4,000 kg.

[0396] Clause 6. The wing assembly of any preceding clause, wherein the ratio of the mass of the rechargeable battery to the mass of the wing structure is at least 4.5, at least 5, at least 5.5, at least 6, or at least 6.5.

[0397] Clause 7. The wing structure is adapted to be attached to the underside of the fuselage; and / or 10. The wing assembly of any preceding clause, wherein the wing structure comprises a foldable tip.

[0398] Article 8. A maximum take-off mass (MTOM) of at least 8,618 kg; an electric range factor (ERF) of at least 6; 1. A wing assembly as described in any preceding clause for a wing-and-body design aircraft having

[0399] Clause 9. An aircraft comprising a wing assembly according to any one of clauses 1 to 8, A plurality of thrusters; a primary energy source configured to power the propulsors, the primary energy source comprising a rechargeable battery; Furthermore, This aircraft is A maximum take-off mass (MTOM) of at least 8,618 kg; an electric range factor (ERF) of at least 6; An aircraft having:

[0400] Clause 10. An aircraft as described in clause 9 having an ERF of at least 6.5, at least 7, at least 7.5, at least 8, at least 8.5, or at least 9.

[0401] Clause 11. An aircraft as referred to in clause 9 or 10 having an MTOM of at least 20,000 kg, at least 40,000 kg, at least 60,000 kg, at least 80,000 kg or at least 100,000 kg.

[0402] Clause 12. An aircraft as described in any of clauses 9 to 11 having a lift to drag ratio "L / D" of at least 15, at least 17, at least 19, or at least 21.

[0403] Clause 13. An aircraft as claimed in any of clauses 9 to 12, having a wing volume to fuselage volume ratio "WV / FV" of at least 0.24, at least 0.27, at least 0.30 or at least 0.33, or a wing area to fuselage area ratio of at least 0.5, at least 0.7, at least 0.9, at least 1.1 or at least 1.3.

[0404] Clause 14. The vessel further comprises a secondary energy source configured to power the thruster, the secondary energy source having a higher energy density than the primary energy source; Optionally, the secondary energy source comprises a liquid fuel, a non-fossil liquid fuel, a non-rechargeable battery, a metal-air battery, or an aluminum-air battery.

[0405] Clause 15. An aircraft as described in any one of clauses 9 to 14 or a wing assembly as described in any one of clauses 1 to 8, wherein the rechargeable battery is configured to be recharged in situ.

[0406] The following numbered examples are also described herein:

[0407] Example 1. An aircraft having a fuselage and a wing, said wing and said fuselage being sized and configured to give the aircraft a maximum take-off mass (MTOM) of between 20 and 150 tons and a wing volume to fuselage volume ratio (WV / FV) of between 0.20 and 0.50.

[0408] Example 2. An aircraft as described in Example 1 having an MTOM of 50 to 100 tons.

[0409] Example 3. The aircraft of Example 1 or 2 having a WV / FV of 0.25 to 0.40.

[0410] Example 4. The aircraft of any one of the preceding examples, having a lift-to-drag ratio (L / D) of 20-30.

[0411] Example 5. The aircraft of any one of the preceding examples having a wing-body design.

[0412] Example 6. An aircraft as described in any one of the preceding examples that is in the transport aircraft category as defined by Title 14 of the Code of Federal Regulations (14 CFR).

[0413] Example 7. An aircraft as described in any one of the preceding examples that is certified in accordance with EASA CS-25 or FAA Part 25, or any special conditions applicable to electric aircraft.

[0414] Example 8. The aircraft of any one of the preceding examples, using battery power for propulsion.

[0415] Example 9. The aircraft of example 8, wherein at least 80% of the batteries used for propulsion are installed in the wings.

[0416] Example 10. The aircraft of example 9, wherein all or substantially all of the batteries used for propulsion are located in the wings.

[0417] Example 11. The aircraft of any one of the preceding examples having one wing.

[0418] Example 12. The aircraft of example 11, wherein the wings are attached to the fuselage at the lower fuselage half.

[0419] Example 13. The aircraft of example 11, wherein the wings are attached to or near the bottom of the fuselage.

[0420] Example 14. The aircraft of any one of the preceding examples, having a seating capacity of at least 70.

[0421] Example 15. The aircraft of Example 14 having a seating capacity of at least 80.

[0422] Example 16. The aircraft of Example 15 having a seating capacity of approximately 90.

[0423] Example 17. The aircraft of any one of Examples 8 to 16, wherein the battery has an energy / weight ratio of at least 240 Wh / kg, preferably 360 Wh / kg, more preferably 440 Wh / kg.

[0424] Example 18. An aircraft described in any one of the preceding examples having a wingspan in flight of more than 35 meters.

[0425] Example 19. The aircraft of Example 18, having a wingspan of approximately 42 meters in flight.

[0426] Example 20. The aircraft of example 19 having foldable wingtips.

[0427] Example 21. The aircraft of any one of Examples 8 to 20, having a plurality of propellers driven by electric motors.

[0428] Example 22. The aircraft of example 21 having at least six propellers.

[0429] Example 23. The aircraft of example 22 having at least eight propellers.

[0430] Example 24. The aircraft of any one of Examples 8 to 23, having a battery-powered range of at least 500 km.

[0431] Example 25. The aircraft of example 24 having a battery-powered range of at least 800 km.

[0432] Example 26. The aircraft of example 25 having a battery-powered range of at least 1000 km.

[0433] Example 27. The aircraft of any one of Examples 8 to 26, having a range extender for providing power to supplement battery power.

[0434] Example 28. The aircraft of example 27, wherein the range extender is capable of operating on liquid fuel.

[0435] Example 29. The aircraft of example 28, wherein the liquid fuel is a non-fossil fuel.

[0436] Example 30. A wing for use on an aircraft according to any one of the preceding examples.

[0437] Example 31. A range extender for use with an aircraft according to any one of the preceding examples.

[0438] Example 32. A battery for use in an aircraft according to any one of the preceding examples.

Claims

1. An aircraft, a fuselage and at least one wing arranged in a wing-body design; A plurality of thrusters; a primary energy source configured to power the propulsors, the primary energy source comprising a rechargeable battery; and An aircraft comprising: The aircraft a maximum take-off mass (MTOM) of at least 8,618 kg; an electric range factor (ERF) of at least 6; An aircraft having:

2. 10. The aircraft of claim 1, having an ERF of at least 6.5, at least 7, at least 7.5, at least 8, at least 8.5, or at least 9.

3. 3. The aircraft of claim 1 or 2, having a MTOM of at least 20,000 kg, at least 40,000 kg, at least 60,000 kg, at least 80,000 kg, or at least 100,000 kg.

4. 10. An aircraft as claimed in any preceding claim, having a Maximum Payload Mass "PLM" of at least 3,000 kg, at least 6,000 kg, at least 9,000 kg, at least 12,000 kg or at least 15,000 kg.

5. 10. The aircraft of any preceding claim, having at least 50%, at least 60%, at least 70%, at least 80%, at least 90%, or at least 95% of its total mass associated with the at least one wing being rechargeable batteries.

6. 10. An aircraft as claimed in any preceding claim, having a lift to drag ratio "L / D" of at least 15, at least 17, at least 19, or at least 21.

7. a secondary energy source configured to power the propulsors; the secondary energy source has a greater energy density than the primary energy source; Optionally, the secondary energy source comprises a liquid fuel, a non-fossil liquid fuel, a non-rechargeable battery, a metal-air battery, or an aluminium-air battery.

8. 10. An aircraft as claimed in any preceding claim, having a wing volume to fuselage volume ratio "WV / FV" of at least 0.20, at least 0.24, at least 0.25, at least 0.27, at least 0.30, or at least 0.

33.

9. 10. The aircraft of any preceding claim, having a wing wetted area to fuselage wetted area ratio of at least 0.50, optionally at least 0.55, optionally at least 0.60, optionally at least 0.65, optionally at least 0.70, optionally at least 0.75, optionally at least 0.80, optionally at least 0.85, optionally at least 0.90, optionally at least 0.95, optionally at least 1.00, optionally at least 1.10, optionally at least 1.20, or optionally at least 1.

30.

10. A wing loading smaller than WL2 or within the range of WL1 to WL2, said WL2 is 850 kg / m, or optionally 800 kg / m, or optionally 750 kg / m, or optionally 700 kg / m, or optionally 650 kg / m, or optionally 600 kg / m, or optionally 550 kg / m; 8. The aircraft of any preceding claim, wherein WL1 is 100 kg / m, or optionally 150 kg / m, or optionally 200 kg / m, or optionally 250 kg / m, or optionally 300 kg / m, or optionally 350 kg / m, or optionally 400 kg / m.

11. the at least one wing is attached to the underside of the fuselage; and / or the at least one wing having a foldable tip; 10. An aircraft according to any preceding claim.

12. 1. A wing assembly for a wing-and-body design aircraft, said wing assembly comprising: The wing structure and a primary energy source configured to power a plurality of propulsors, the primary energy source comprising a rechargeable battery; Equipped with 10. A wing assembly, wherein the wing structure has a mass of at least 600 kg, and wherein the ratio of the mass of the rechargeable battery to the mass of the wing structure is at least 4.

13. the wing structure has a mass of at least 1,000 kg, at least 2,000 kg, at least 3,000 kg, or at least 4,000 kg; and / or 13. The wing assembly of claim 12, wherein the ratio of the mass of the rechargeable battery to the mass of the wing structure is at least 4.5, at least 5, at least 5.5, at least 6, or at least 6.

5.

14. the wing structure is configured to be attached to the underside of a fuselage; and / or the wing structure comprises a foldable tip; A wing assembly according to claim 12 or 13.

15. a maximum take-off mass (MTOM) of at least 8,618 kg; an electric range factor (ERF) of at least 6; A wing assembly according to any one of claims 12 to 14 for a wing-and-body design aircraft having

16. An aircraft according to any one of claims 1 to 11, comprising a wing assembly according to any one of claims 12 to 15.

17. the plurality of thrusters comprises 6, 8, 10, 12 or 14 thrusters; and / or the rechargeable battery is configured to be recharged in situ; An aircraft according to any one of claims 1 to 11 or 16 or a wing assembly according to any one of claims 12 to 15.

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