System and method for flight control of EVTOL aircraft

Abstract

Aspects of the present disclosure relate to flight control of electric aircraft and other vehicles. In one embodiment, an aircraft is disclosed which comprises a fuselage, one or more flight control computers configured to provide control signals, four sets of electrically driven propellers, the first set being positioned in front of the second set and both on one side of the fuselage, the third set being positioned in front of the fourth set and both on the other side of the fuselage, and a plurality of electric buses, the one or more flight control computers configured to provide control signals to at least one of the first set of propellers, at least one of the fourth set of propellers, and at least one of the third set of propellers and at least one of the second set of propellers via one of the plurality of electric buses.

Description

【Technical Field】 【0001】 Cross - Reference to Related Applications This application claims priority to U.S. Application No. 18 / 468,421, filed on September 15, 2023, titled "SYSTEMS AND METHODS FOR FLIGHT CONTROL OF EVTOL AIRCRAFT" (Attorney Docket No. 16499.0011 - 00000), which claims priority to U.S. Provisional Application No. 63 / 504,958, filed on May 30, 2023, titled "SYSTEMS AND METHODS FOR FLIGHT CONTROL OF EVTOL AIRCRAFT" (Attorney Docket No. 16499.6005 - 00000). The disclosures of all of these are hereby incorporated by reference in their entirety for all purposes. 【0002】 This disclosure generally relates to powered aircraft. More specifically, and without limitation, this disclosure relates to technological innovations in aircraft driven by electric propulsion systems. Certain aspects of this disclosure generally relate to flight control in aircraft and other types of vehicles driven by electric propulsion systems, as well as systems and methods for flight control of aircraft in flight simulators and video games. Other aspects of this disclosure generally relate to improvements in flight control systems and methods that provide certain advantages to aerial vehicles and can be used in other types of vehicles. 【Summary of the Invention】 【0003】 Aspects of this disclosure relate to flight control of electric aircraft and other vehicles. In one embodiment, an aircraft is disclosed which comprises a fuselage, one or more flight control computers configured to provide control signals, a first set of electric-driven propellers and a second set of electric-driven propellers disposed on one side of the fuselage, the first set comprising the first set of electric-driven propellers and the second set of electric-driven propellers disposed in front of the second set, and a third set of electric-driven propellers and a fourth set of electric-driven propellers disposed on the other side of the fuselage, the third set comprising the third set of electric-driven propellers and the fourth set of electric-driven propellers disposed in front of the fourth set. The system comprises four sets of electrically driven propellers and a plurality of electric buses coupled to one or more flight control computers, wherein one or more flight control computers are configured to provide control signals to at least one of the first set of propellers and at least one of the fourth set of propellers via one of the plurality of electric buses, and one or more flight control computers are configured to provide control signals to at least one of the third set of propellers and at least one of the second set of propellers via another of the plurality of electric buses. [Brief explanation of the drawing] 【0004】 [Figure 1] An exemplary VTOL aircraft consistent with the disclosed embodiments is shown. [Figure 2] An exemplary VTOL aircraft consistent with the disclosed embodiments is shown. [Figure 3] An exemplary top view of a VTOL aircraft, consistent with the disclosed embodiments, is shown. [Figure 4] An exemplary propeller rotation of a VTOL aircraft, consistent with the disclosed embodiments, is illustrated. [Figure 5] This shows an exemplary power connection in a VTOL aircraft, consistent with the disclosed embodiments. [Figure 6] An exemplary architecture of an electric propulsion unit, consistent with the disclosed embodiments, is shown. [Figure 7] An exemplary top view of a VTOL aircraft, consistent with the disclosed embodiments, is shown. [Figure 8A] An exemplary flight control signaling architecture consistent with the disclosed embodiments is shown. [Figure 8B] An exemplary flight control signaling architecture consistent with the disclosed embodiments is shown. [Figure 8C] An exemplary flight control signaling architecture consistent with the disclosed embodiments is shown. [Figure 8D] An exemplary flight control signaling architecture consistent with the disclosed embodiments is shown. [Figure 9A] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 9B] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 9C] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 9D] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 10A] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 10B] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 10C] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 10D]Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 11A] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 11B] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 11C] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 11D] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 12A] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 12B] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 12C] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 12D] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 13A] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 13B] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 13C] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Figure 13D] Exemplary flight control signaling architectures and their components, consistent with the disclosed embodiments, are illustrated and described. [Modes for carrying out the invention] 【0005】 This disclosure primarily describes systems, components, and technologies for use in aircraft. Aircraft may be piloted aircraft, unpiloted aircraft (e.g., UAVs), drones, helicopters, and / or airplanes. An aircraft comprises a physical body and one or more components (e.g., wings, tails, propellers) configured to enable the aircraft to fly. An aircraft may include any configuration including at least one propeller. In some embodiments, an aircraft is driven by one or more electric propulsion systems (hereinafter referred to as electric propulsion units or "EPUs"). Aircraft may be fully electric, hybrid, or gas-powered. For example, in some embodiments, an aircraft is a tiltrotor configured for frequent (e.g., more than 50 flights per working day), short-duration flights (e.g., less than 100 miles per flight) over, into, and outside densely populated areas. An aircraft may be configured to carry 4 to 6 passengers or commuters expecting a comfortable experience with low noise and low vibration. Therefore, it may be desirable that aircraft components be configured and designed to withstand frequent use without wear and tear, generate less heat and vibration, and that the aircraft include mechanisms to effectively control and manage the heat or vibration generated by the components. Furthermore, some of these aircraft may be intended to operate in close proximity to each other over congested metropolitan areas. Therefore, it may be desirable that their components be configured and designed to generate low levels of noise inside and outside the aircraft and have various safety and backup mechanisms. For example, for safety reasons, it may be desirable that the aircraft be propelled by a distributed propulsion system to avoid the risk of a single point of failure and to be able to perform conventional takeoffs and landings on runways. Furthermore, it may be desirable that the aircraft be able to safely take off and land vertically from or into relatively small or limited spaces (e.g., vertiports, parking lots, or private roads) compared to conventional airport runways, while transporting several passengers or commuters along with their baggage.These usage requirements can impose design constraints on the size, weight, and operating efficiency (e.g., drag, energy usage) of the aircraft, and the design constraints can affect the design and configuration of aircraft components. 【0006】 The disclosed embodiments provide a novel and improved configuration of aircraft components and / or specified design criteria different from those of conventional aircraft components that are not found in conventional aircraft. Such alternative configurations and design criteria are combined to address the drawbacks and problems inherent in conventional components, resulting in the various configurations and designs of components for an aircraft (e.g., an electric aircraft or a hybrid electric aircraft) driven by an electric propulsion system obtained in the embodiments disclosed herein. 【0007】 In some embodiments, an aircraft driven by an electric propulsion system can be designed to be capable of both vertical and conventional takeoff and landing, with a distributed electric propulsion system that enables vertical flight, horizontal and lateral flight, and transitions. Thrust can be generated by supplying high-voltage power to a plurality of electric engines of the distributed electric propulsion system, which can include components necessary to convert the high-voltage power into mechanical shaft power to rotate the propellers. The embodiments disclosed herein can involve optimizing the energy density of the electric propulsion system. Embodiments can include an electric engine connected to an on-board power source, which can include a device capable of storing energy such as a battery or a capacitor, and one or more systems for utilizing or generating electricity such as a fuel-powered generator or a solar panel array. Some of the disclosed embodiments provide a reduction in the weight and space of components within the aircraft to increase the efficiency and performance of the aircraft. The disclosed embodiments also improve safety in passenger transportation by using new and improved safety protocols and system redundancy in case of failure to minimize any single point of failure in the aircraft propulsion system. 【0008】 In some embodiments, the aircraft may include one or more wings and / or wing structures (hereinafter interchangeably referred to as wing structures or wings). As described herein, wing structures may include conventional wings, empennages, stabilizers, winglets, or any other wing that may be designed to provide at least some lift. It should be understood that any description using the term wing may be equivalent to using any other type of wing structure. 【0009】 In some embodiments, the distributed electric propulsion system may include twelve electric engines that may be mounted on booms in front of and behind the main wing of the aircraft. A subset of the electric engines, such as those mounted in front of the main wing, may be tiltable during flight between a horizontally oriented attitude (e.g., generating forward thrust for cruising) and a vertically oriented attitude (e.g., generating vertical lift for takeoff, landing, and hovering). For example, the vertically oriented attitude may include a propeller facing upward or a propeller facing downward. The propellers of the forward electric engines may rotate in a clockwise or counterclockwise direction. The propellers may rotate in a direction opposite to that of adjacent propellers. The rear electric engines may be fixed in a vertically oriented attitude (e.g., generating vertical lift). The propellers may also rotate in a clockwise or counterclockwise direction. In some embodiments, the difference in the rotation direction may be achieved using the engine rotation direction. In other embodiments, the engines may all rotate in the same direction, and gearing may be used to achieve different propeller rotation directions. The vertical lift is defined as the lift in the vertical direction configured within a range of 90 degrees ± 15%. The horizontal thrust is defined as the thrust in the horizontal direction consisting of a range of 0 degrees ± 15%. 【0010】 In some embodiments, an aircraft may have a number of electric engines in various combinations of forward-engine and rear-engine configurations. For example, an aircraft may have any other combination of forward-engine and rear-engine engines, including six forward-engine and six rear-engine engines, five forward-engine and five rear-engine engines, four forward-engine and four rear-engine engines, three forward-engine and three rear-engine engines, two forward-engine and two rear-engine engines, or embodiments in which the number of forward-engine and rear-engine engines are not equal. 【0011】 In some embodiments, any number or combination of the aircraft's electric engines may be configured to be tiltable. For example, an aircraft may have any other number or combination of forward and rear engines, including embodiments in which the number of forward electric engines and rear electric engines are not equal, or in which the combination of tiltable forward electric engines and tiltable rear electric engines is not equal. 【0012】 In some embodiments, for vertical take-off and landing (VTOL) missions, forward and rear electric engines may provide vertical thrust during take-off and landing. During the forward flight phase, the forward electric engine may provide horizontal thrust, while the rear electric engine's propeller may be retracted to a fixed position to minimize drag. The rear electric engine may be actively retracted while maintaining position monitoring. Transitions from vertical to horizontal flight and vice versa may be achieved via a tilt propeller subsystem. The tilt propeller subsystem may redirect thrust, primarily vertical, during vertical flight mode to a horizontal or nearly horizontal direction during the forward flight cruising phase. A variable pitch mechanism may change the collective angle of the forward electric engine's propeller hub assembly blades for operation during the hovering, transition, and cruising phases. 【0013】 In some embodiments, in conventional take-off and landing (CTOL) missions, the forward electric engines may provide horizontal thrust for fixed-wing take-off, cruising, and landing, while the wings may provide vertical lift. In some embodiments, the rear electric engines may not be used to generate thrust during CTOL missions, and the rear propellers may be retracted into a fixed position. In other embodiments, the rear electric engines may be used at reduced power to shorten the length of CTOL take-off or landing. 【0014】 The disclosed embodiments provide systems, subsystems, and components for novel VTOL aircraft having various combinations of electric propulsion and cooling systems that maximize performance while minimizing weight. For example, the novel VTOL aircraft may be referred to herein as an eVTOL (electric vertical take-off and landing) aircraft. 【0015】 In some embodiments, the electric propulsion systems described herein may generate thrust by supplying high-voltage (HV) power to electric engines, which in turn convert the HV power into mechanical shaft power used to rotate a propeller. The aircraft described herein may include a plurality of electric engines mounted forward and backward on the wings. The engines may be mounted directly on the wings or on one or more booms attached to the wings. The amount of thrust generated by each electric engine may be controlled by torque commands from a flight control system (FCS) via a digital communication interface to each electric engine. Embodiments may include forward-rotating electric engines whose orientation or inclination can be changed. Some embodiments include forward-rotating engines which may be of clockwise (CW) or counterclockwise (CCW) type. The forward electric propulsion subsystem may consist of a multi-blade adjustable-pitch propeller, as well as a variable-pitch subsystem. 【0016】 In some embodiments, the aircraft may include a rear electric engine or lifter, which can be of the clockwise (CW) or counterclockwise (CCW) type. Some embodiments may include a rear electric engine utilizing a multi-blade fixed-pitch propeller. 【0017】 As described herein, the orientation and use of electric propulsion system components may change throughout the operation of the aircraft. In some embodiments, during vertical takeoff and landing, the forward and rearward propulsion systems may provide vertical thrust during takeoff and landing. During the flight phase when the aircraft is in forward flight mode, the forward propulsion system may provide horizontal thrust, while the propellers of the rearward propulsion system may be retracted to a fixed position to minimize drag. The rearward electric propulsion system may be actively retracted with position monitoring. Some embodiments may include transitions from vertical to horizontal flight and vice versa. In some embodiments, the transition may be achieved via a tilt propeller system (TPS). The TPS redirects the thrust, which is primarily vertical during vertical flight mode, to primarily horizontal during forward flight mode. Some embodiments may include a variable pitch mechanism that can change the collective angle of the propeller blades of the forward propulsion system for operation during the hovering phase, cruising phase, and transition phase. Some embodiments may include a conventional takeoff and landing (CTOL) configuration such that the tilter provides horizontal thrust for the fixed-wing takeoff phase, cruising phase, and landing phase. In some embodiments, the rear electric engine does not need to be used to generate thrust during CTOL missions, and the rear propeller is retracted into a fixed position to minimize drag. 【0018】 Herein, exemplary embodiments are given in detail, and examples are illustrated in the accompanying drawings. The following description is given with reference to the accompanying drawings, and unless otherwise specified, the same or similar reference numerals in different drawings represent the same or similar elements. The implementations described below in the exemplary embodiments do not represent all implementations consistent with the present disclosure. Rather, they are merely examples of apparatus and methods consistent with embodiments relating to the subject matter described in the accompanying claims. 【0019】 Figure 1 is an illustrative perspective view of an exemplary VTOL aircraft consistent with the disclosed embodiments. Figure 2 is another illustrative perspective view of an exemplary VTOL aircraft in an alternative configuration consistent with embodiments of the present disclosure. Figures 1 and 2 illustrate VTOL aircraft 100, 200 in a cruising configuration and a vertical takeoff, landing, and hovering configuration (also referred to herein as a “lift” configuration), respectively, consistent with embodiments of the present disclosure. Elements corresponding to Figures 1 and 2 may have similar figures and may refer to similar elements of aircraft 100, 200. Aircraft 100, 200 may include fuselages 102, 202, wings 104, 204 mounted on the fuselages 102, 202, and one or more rear stabilizers 106, 206 mounted on the rear of the fuselages 102, 202. Multiple lift propellers 112, 212 may be mounted on wings 104, 204 and configured to provide lift for vertical takeoff, landing, and hovering. Multiple tilt propellers 114, 214 may be mounted on wings 104, 204 and may be tiltable between a lift configuration, as shown in Figure 2, which provides a portion of the lift required for vertical takeoff, landing, and hovering, and a cruising configuration, as shown in Figure 1, which provides forward thrust to the aircraft 100 for horizontal flight. As used herein, the lift configuration of a tilt propeller refers to any tilt propeller orientation in which the tilt propeller thrust primarily provides lift to the aircraft, and the cruising configuration of a tilt propeller refers to any tilt propeller orientation in which the tilt propeller thrust primarily provides forward thrust to the aircraft. 【0020】 In some embodiments, the lift propellers 112, 212 may be configured to provide only lift, with all horizontal thrust provided by the tilt propellers. Thus, the lift propellers 112, 212 may be configured in a fixed position and may generate thrust only during the takeoff, landing, and hovering phases of flight. On the other hand, the tilt propellers 114, 214 may be tilted upward in a lift configuration in which the thrust from the propellers 114, 214 is directed downward to provide additional lift. 【0021】 For forward flight, the tilt propellers 114, 214 can be tilted from a lift configuration to a cruising configuration. In other words, the orientation of the tilt propellers 114, 214 can change from an orientation in which the thrust of the tilt propellers is directed downward (to provide lift during vertical takeoff, landing, and hovering) to an orientation in which the thrust of the tilt propellers is directed aft (to provide forward thrust to the aircraft 100, 200). The tilt propeller assembly for a particular electric engine can be tilted around an axis of rotation defined by the mounting point connecting the boom and the electric engine. When the aircraft 100, 200 is in full forward flight, lift can be provided entirely by the wings 104, 204. On the other hand, in the cruising configuration, the lift propellers 112, 212 can be shut off. The blades 120, 220 of the lift propellers 112, 212 can be held in a low-drag position for aircraft cruising. In some embodiments, each of the lift propellers 112, 212 may have two blades 120, 220 that can be locked for cruising in a minimum-drag position where one blade is directly in front of the other, as illustrated in Figure 1. In some embodiments, the lift propellers 112, 212 may have three or more blades. In some embodiments, the tilt propellers 114, 214 may include more blades 116, 216 than the lift propellers 112, 212. For example, as illustrated in Figures 1 and 2, each of the lift propellers 112, 212 may include, for example, two blades, while each of the tilt propellers 114, 214 may include more blades, such as the five blades shown. In some embodiments, each of the tilt propellers 114, 214 may have 2 to 5 blades, and possibly more, depending on the design considerations and requirements of the aircraft. 【0022】 In some embodiments, the aircraft may include a single wing 104, 204 on each side of the fuselage 102, 202 (or a single wing extending over the entire aircraft). At least some of the lift propellers 112, 212 may be located behind the wings 104, 204, and at least some of the tilt propellers 114, 214 may be located in front of the wings 104, 204. In some embodiments, all of the lift propellers 112, 212 may be located behind the wings 104, 204, and all of the tilt propellers 114, 214 may be located in front of the wings 104, 204. According to some embodiments, all of the lift propellers 112, 212 and tilt propellers 114, 214 may be mounted on the wings, i.e., the lift propellers or tilt propellers may not be mounted on the fuselage. In some embodiments, the lift propellers 112, 212 may all be located behind the wings 104, 204, and the tilt propellers 114, 214 may all be located in front of the wings 104, 204. According to some embodiments, all the lift propellers 112, 212 and the tilt propellers 114, 214 may be located inside the ends of the wings 104, 204. 【0023】 In some embodiments, lift propellers 112, 212 and tilt propellers 114, 214 may be mounted on the wings 104, 204 by booms 122, 222. Booms 122, 222 may be mounted below, above, and / or incorporated into the wing profile. In some embodiments, lift propellers 112, 212 and tilt propellers 114, 214 may be mounted directly on the wings 104, 204. In some embodiments, one lift propeller 112, 212 and one tilt propeller 114, 214 may be mounted on each boom 122, 222. Lift propellers 112, 212 may be mounted at the rear end of booms 122, 222, and tilt propellers 114, 214 may be mounted at the front end of booms 122, 222. In some embodiments, lift propellers 112, 212 may be mounted in fixed positions on booms 122, 222. In some embodiments, tilt propellers 114, 214 may be mounted to the front end of booms 122, 222 via hinges. The tilt propellers 114, 214 may be mounted on booms 122, 222 such that when the tilt propellers 114, 214 are in a cruising configuration, they align with the body of booms 122, 222, forming a continuous extension of the front end of booms 122, 222 that minimizes drag for forward flight. 【0024】 In some embodiments, the aircraft 100, 200 may include, for example, one wing on each side of the fuselage 102, 202, or a single wing extending across the entire aircraft. According to some embodiments, at least one wing 104, 204 is a high wing mounted on the upper side of the fuselage 102, 202. According to some embodiments, the wing includes control surfaces such as flaps and / or ailerons. According to some embodiments, the wings 104, 204 may be designed with a profile that reduces drag during forward flight. In some embodiments, the wingtip profile may be curved and / or tapered to minimize drag. 【0025】 In some embodiments, the rear stabilizers 106, 206 include control surfaces such as one or more rudders, one or more elevators, and / or one or more combined rudder-elevators. The wings(s) may have any preferred design. In some embodiments, the wings have tapered leading edges. 【0026】 In some embodiments, a lift propeller 112, 212 or tilt propeller 114, 214 may tilt relative to at least one other lift propeller 112, 212 or tilt propeller 114, 214. As used herein, canting refers to the relative orientation of the axis of rotation of a lift propeller / tilt propeller around a line parallel to the longitudinal direction, similar to the roll degrees of freedom of an aircraft. The tilting of a lift propeller and / or tilt propeller may help minimize damage from propeller rupture by oriented the plane of rotation of the lift propeller / tilt propeller disc (the blades and the hub on which those blades are mounted) so as not to intersect with critical parts of the aircraft (such as areas of the fuselage where a person may be positioned, critical flight control systems, batteries, adjacent propellers, etc.) or other propeller discs, and may provide enhanced yaw control during flight. 【0027】 In some embodiments, the lift propellers 112, 212 and the tilt propellers 114, 214 can be directly mounted on the non-wing elements of the aircraft 100, 200. For example, in some embodiments, the lift propellers 112, 212 and the tilt propellers 114, 214 can be directly mounted on the fuselage 102, 202. Furthermore, in some embodiments, the lift propellers 112, 212 and the tilt propellers 114, 214 can be directly mounted on the rear stabilizers 106, 206. 【0028】 Figure 3 is an illustrative top view of an exemplary VTOL aircraft consistent with embodiments of the present disclosure. The aircraft 300 shown in the figure may be a top view of aircraft 100, 200 shown in Figures 1 and 2, respectively. As considered herein, the aircraft 300 may include twelve electric propulsion systems spread out over the aircraft 300. In some embodiments, the distribution of electric propulsion systems may include six forward electric propulsion systems 314 and six rear electric propulsion systems 312 mounted on the forward and rear booms of the main wing 304 of the aircraft 300. In some embodiments, the length of the trailing end of the boom 324 from the wing 304 to the lift propeller 312 may include similar trailing ends of the boom 324 across a number of trailing ends of the boom. In some embodiments, the length of the trailing end of the boom may vary across six exemplary trailing ends of the boom. For example, each trailing end of the boom 324 may include a different length from the wing 304 to the lift propeller 312, or a subset of the trailing ends of the boom may have similar lengths. In some embodiments, the front end of the boom 322 may include varying lengths from the wings 304 to the tilt propeller 314 across the entire front end of the boom. For example, as shown in Figure 3, the length of the front end of the boom 322 from the tilt propeller 314 closest to the fuselage to the wings 304 may include a length longer than the length of the front end of the boom 322 from the wings 304 to the tilt propeller 314 furthest from the fuselage. Some embodiments may include a boom front end having similar lengths across the entire six exemplary front ends of the boom, or any other length distribution, from the wings 304 to the tilt propeller 314. Some embodiments may include an aircraft 300 having eight electric propulsion systems, with four forward electric propulsion systems 314 and four rear electric propulsion systems 312, or any other distribution of forward and rear electric propulsion systems, including embodiments in which the number of forward electric propulsion systems 314 is less than or greater than the number of rear electric propulsion systems 312. Furthermore, Figure 3 illustrates an exemplary embodiment of a VTOL aircraft 300, which has a forward propeller 314 oriented horizontally for horizontal flight and a rear propeller blade 320 in a retracted position for forward flight. 【0029】 As disclosed herein, forward and rear electric propulsion systems may be of the clockwise (CW) or counterclockwise (CCW) type. Some embodiments may include a variety of forward electric propulsion systems having a mixture of both CW and CCW types. In some embodiments, the rear electric propulsion system may have a mixture of CW and CCW type systems among the rear electric propulsion systems. 【0030】 Figure 4 is a schematic diagram illustrating exemplary propeller rotation of a VTOL aircraft consistent with the disclosed embodiments. The aircraft 400 shown in the figure may be a top view of aircraft 100, 200, and 300 shown in Figures 1, 2, and 3, respectively. In aircraft 400, three of the forward electric propulsion systems are CW type 424, and the remaining three forward electric propulsion systems are CCW type 426 This may include six forward electric propulsion systems. In some embodiments, three rearward electric propulsion systems may be CCW type 428, and the remaining three rearward electric propulsion systems may be CW type 430. Some embodiments may include an aircraft 400 having four forward electric propulsion systems and four rearward electric propulsion systems, each having two CW type and two CCW type. In some embodiments, propellers may be reversed relative to adjacent propellers to cancel torque steer generated by the rotation of the propellers and experienced by the fuselage or wings of the aircraft. In some embodiments, the difference in rotation direction may be achieved using the engine rotation direction. In other embodiments, the engines may all rotate in the same direction, and gearing may be used to achieve different propeller rotation directions. 【0031】 Some embodiments may include an aircraft 400 having a forward electric propulsion system and a rear electric propulsion system, where the quantities of CW type 424 and CCW type 426 are not equal between forward electric propulsion systems, between rear electric propulsion systems, or between forward electric propulsion systems and rear electric propulsion systems. 【0032】 Figure 5 is a schematic diagram illustrating exemplary power connections in a VTOL aircraft, consistent with the disclosed embodiments. A VTOL aircraft may have various power systems connected to diagonally opposed electric propulsion systems. In some embodiments, the power systems may include high-voltage power systems. In some embodiments, high-voltage power systems may be connected to electric engines via high-voltage channels. In some embodiments, the aircraft 500 may include six power systems, including batteries 526, 528, 530, 532, 534, and 536 housed within the wings 570 of the aircraft 500. In some embodiments, the aircraft 500 may include six forward electric propulsion systems having six electric engines 502, 504, 506, 508, 510, and 512, and six rearward electric propulsion systems having six electric engines 514, 516, 518, 520, 522, and 524. In some embodiments, batteries may be connected to diagonally opposed electric engines. In such a configuration, the first power system 526 may supply power to the electric engine 502 via a power connection channel 538 and to the electric engine 524 via a power connection channel 540. In some embodiments, the first power system 526 may be paired with a fourth power system 532 via a power connection channel 542 having a fuse to prevent excessive current from flowing through power systems 526 and 532. In addition to this embodiment, the VTOL aircraft 500 may include a second power system 528 paired with a fifth power system 534 via a power connection channel 548 having a fuse, which may supply power to electric engines 510 and 516 via power connection channels 544 and 546, respectively. In some embodiments, a third power system 530 may be paired with a sixth power system 536 via a power connection channel 554 having a fuse, which may supply power to electric engines 506 and 520 via power connection channels 550 and 552, respectively. The fourth power system 532 may also supply power to the electric engines 508 and 518 via power connection channels 556 and 558, respectively. The fifth power system 534 may also supply power to the electric engines 504 and 522 via power connection channels 560 and 562, respectively.The sixth power grid 536 may also supply power to the electric engines 512 and 514 via power connection channels 564 and 566, respectively. 【0033】 As disclosed herein, an electric propulsion system may include an electric engine connected to a high-voltage power system, such as a battery located within the aircraft, via a high-voltage channel or power connection channel. Some embodiments may include various batteries housed within the aircraft wing, having high-voltage channels that lead to the electric propulsion system throughout the aircraft, including the wings and boom. In some embodiments, multiple high-voltage power systems may be used to create an electric propulsion system with multiple high-voltage power sources to avoid the risk of a single point of failure. In some embodiments, the aircraft may include multiple electric propulsion systems that can be pattern-wired to various batteries or power sources housed throughout the aircraft. It is recognized that such a configuration may be beneficial in avoiding the risk of a single point of failure, where the failure of one battery or power source could result in a portion of the aircraft being unable to maintain the amount of thrust necessary to continue flight or perform a controlled landing. For example, if a VTOL has two forward electric propulsion systems and two rearward electric propulsion systems, the forward and rearward electric propulsion systems on both sides of the VTOL aircraft may be connected to the same high-voltage power system. In such a configuration, if one high-voltage power system fails, one of the forward and one of the rear electric propulsion systems located on opposite sides of the VTOL aircraft may remain operational, potentially providing a more balanced flight or landing compared to forward and rear electric propulsion systems failing on the same side of the VTOL aircraft. Some embodiments may include four forward and four rear electric propulsion systems in which diagonally opposed electric engines are connected to a common battery or power source. Some embodiments may include various configurations of electric engines electrically connected to a high-voltage power system so that the risk of a single point of failure in the event of a power failure can be avoided, allowing the flight phase in which the failure occurs to continue or allowing the aircraft to perform an alternative flight phase in response to the failure. 【0034】 As discussed above, an electric propulsion system may include an electric engine that provides mechanical shaft power to a propeller assembly to generate thrust. In some embodiments, the electric engine of an electric propulsion system may include a high-voltage power system that supplies high-voltage power to the electric engine and / or a low-voltage system that supplies low-voltage DC power to the electric engine. Some embodiments may include an electric engine(s) that digitally communicates with a flight control system ("FCS") which has one or more flight control computers ("FCC") that can send and receive signals to and from the electric engine, including command and response data or status. Some embodiments may include an electric engine that can receive operating parameters from the FCC, including speed, voltage, current, torque, temperature, vibration, propeller position, and any other values ​​of operating parameters, and communicate the operating parameters to the FCC. 【0035】 In some embodiments, the flight control system may include a system that can communicate with the electric engine to send and receive analog / discrete signals to the electric engine and control a device that can redirect the thrust of the tilt propeller between primarily vertical in vertical flight mode and primarily horizontal in forward flight mode. In some embodiments, this system may be referred to as a tilt propeller system ("TPS") and may be capable of communicating and oriented additional features of the electric propulsion system. 【0036】 Figure 6 illustrates an exemplary architecture and design block diagram of an electric propulsion unit (EPU) 600 consistent with the disclosed embodiments. In some embodiments, the electric propulsion system 602 may include an electric engine subsystem 604 that can supply torque to a propeller subsystem 606 via a shaft to generate thrust for the electric propulsion system 602. In some embodiments, the electric engine subsystem 604 may receive low-voltage DC (LV DC) power from a low-voltage system (LVS) 608. In some embodiments, the electric engine subsystem 604 may receive high-voltage (HV) power from a high-voltage power system (HVPS) 610 that includes at least one battery or other device capable of storing energy. In some embodiments, the high-voltage power system may include two or more batteries or other devices capable of storing energy and supplying high-voltage power to the electric engine subsystem 604. It is recognized that such configurations may be advantageous in that they do not jeopardize a single point of failure where the failure of a single battery would lead to the failure of the electric propulsion system 602. 【0037】 In some embodiments, the electric propulsion system 602 may include an electric engine subsystem 604 that receives signals from and transmits signals to the flight control system 612. In some embodiments, the flight control system 612 may include a flight control computer that can transmit commands to and receive status and data from the electric engine subsystem 604 using Controller Area Network ("CAN") data bus signals. While the CAN data bus signals are used between the flight control computer and the electric engine(s), it should be understood that in some embodiments, it may include any form of communication that has the ability to send and receive data from the flight control computer to and from the electric engine(s). In some embodiments, the flight control system 612 may also include a tilt propeller system ("TPS") 614 that can send and receive analog and discrete data to and from the tilt propeller electric engine subsystem 604. The tilt propeller system 614 may include a device that communicates operating parameters to the electric engine subsystem 604 and articulates the orientation of the propeller subsystem 606, thereby redirecting the thrust of the tilt propeller during various flight phases using mechanical means such as a gearbox assembly, a linear actuator, and any other configuration of components for changing the orientation of the propeller subsystem 606. 【0038】 As will be considered throughout, exemplary VTOL aircraft may have various types of electric propulsion systems, including tilt and lift propellers, including forward-mounted electric engines capable of tilting during various phases of flight, and rear-mounted electric engines that remain in one orientation and may only be active during specific phases of flight (i.e., takeoff, landing, and hovering). 【0039】 In some embodiments, the flight control system may include a system capable of controlling a control surface and actuators associated with the control surface in an exemplary VTOL aircraft. Figure 7 is an illustrative top view of an exemplary VTOL aircraft consistent with embodiments of the present disclosure. The aircraft 700 shown in the figure may be a top view of aircraft 100 and 200 shown in Figures 1 and 2, respectively. In aircraft 700, the control surface may include a flaperon 712 and a rudder-vator 714 in addition to the propeller blades discussed earlier. The flaperon 712 may combine the functions of one or more flaps, one or more ailerons, and / or one or more spoilers. The rudder-vator 714 may combine the functions of one or more rudders and / or one or more elevators. In aircraft 700, the actuators may include a control surface actuator (CSA) associated with the flaperon 712 and the rudder-vator 714 in addition to the electric propulsion system discussed earlier, as will be further discussed below with reference to Figures 8A to 13D. 【0040】 Figures 8A–D illustrate flight control signaling architectures for controlling control surfaces and associated actuators in various embodiments. Figures 7–13D illustrate 12 EPU inverters and associated propeller blades, 6 tilt propeller actuators (TPACs), 6 battery management systems (BMSs), 4 flaperons and associated control surface actuators (CSAs), and 6 rudder swivels and associated CSAs, but aircraft in various embodiments may have any preferred number of these various elements. As shown in Figure 8A, the control surfaces and actuators are arranged as follows: left FCC, lane A (L FCC-A), left FCC, lane B (L FCC-B), right FCC, lane A (R FCC-A), and right FCC, lane BThe aircraft can be controlled by a combination of four flight control computers (FCCs) (R FCC-B), but any other suitable number of FCCs may be used. Each FCC may control all control planes and actuators individually or in any combination of them. In some embodiments, each FCC may include one or more hardware computing processors. In some embodiments, each FCC may utilize a single-threaded or multi-threaded computing process to perform the calculations required to control the control planes and actuators. In some embodiments, all computing processes required to control the control planes and actuators may be performed by a single flight control computer on a single computing thread. 【0041】 An FCC may provide control signals to control surface actuators, including EPU inverters, TPACs, BMSs, flaperon CSAs, and ladder inverter CSAs, via one or more bus systems. For different control surface actuators, the FCC may provide control signals that can be voltage control signals or current control signals, and the control information may be encoded into binary, digital, or analog control signals. In some embodiments, each bus system may be a CAN bus system, e.g., left CAN bus 1, left CAN bus 2, right CAN bus 1, right CAN bus 2, center CAN bus 1, center CAN bus 2 (see Figure 8A). In some embodiments, multiple FCCs may be able to provide control signals via each CAN bus system, and each FCC may be able to provide control signals via multiple CAN bus systems. In the exemplary architecture illustrated in Figure 8A, for example, L FCC-A may provide control signals via left CAN bus 1 and right CAN bus 1, L FCC-B may provide control signals via left CAN bus 1 and central CAN bus 1, R FCC-A may provide control signals via central CAN bus 2 and right CAN bus 2, and R FCC-B may provide control signals via left CAN bus 2 and right CAN bus 2. 【0042】 Furthermore, the flight control signaling architecture may be configured such that, in the event of a failure of any flight control signaling component (e.g., FCC or bus), any loss of control, lift, or thrust with respect to the aircraft's roll, pitch, and / or yaw is substantially symmetric (e.g., with an asymmetry of <±5%, <±10%, <±15%, <±20%, or <±25%), allowing the aircraft to continue stable flight (even if lift or thrust is reduced). For example, referring to Figure 8B, the left CAN bus 1 may provide control signals to inverters 1 and 12, as well as 6 and 7, such that a failure of this bus may result in a substantially symmetric loss of lift or thrust on both sides of the aircraft. For example, a failure of the left CAN bus 1 would result in both EPU 1 and EPU 12 going offline simultaneously, which is likely to result in a substantially symmetric loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the left CAN bus 1 would result in both EPU6 and EPU7 going offline simultaneously, likely resulting in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 1 may provide control signals to TPAC1 and TPAC6 such that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure of the left CAN bus 1 would result in a loss of control of TPAC1 and TPAC6, likely resulting in a symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 1 may provide control signals to the LI flaperon and RI flaperon such that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure of the left CAN bus 1 would result in both the LI flaperon and RI flaperon going offline simultaneously, likely resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 1 may provide control signals to the LO rudder and RO rudder, such that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft.For example, a failure in left CAN bus 1 is likely to result in both the LO rudder and RO rudder going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0043】 Similarly, referring to Figure 8B, the left CAN bus 2 may provide control signals to inverters 3 and 10, as well as 4 and 9, such that a failure of this bus could result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the left CAN bus 2 would result in both EPU 3 and EPU 10 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the left CAN bus 2 would result in both EPU 4 and EPU 9 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 2 may provide control signals to TPAC 3 and TPAC 4, such that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure of the left CAN bus 2 would result in a loss of control of TPAC 3 and TPAC 4, which would likely result in a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. Furthermore, the left CAN bus 2 may provide control signals to the LI and RI flaperons such that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure of the left CAN bus 2 is likely to cause both the LI and RI flaperons to go offline simultaneously, resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 2 may provide control signals to the LO and RO rudder verso such that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure of the left CAN bus 2 is likely to cause both the LO and RO rudder verso to go offline simultaneously, resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. 【0044】 Similarly, referring to Figure 8C, the central CAN bus 1 may provide control signals to inverters 2 and 11, as well as 4 and 9, such that a failure of this bus could result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the central CAN bus 1 would result in both EPU 2 and EPU 11 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the central CAN bus 1 would result in both EPU 4 and EPU 9 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to TPAC 2 and TPAC 4, such that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure of the central CAN bus 1 would result in a loss of control of TPAC 2 and TPAC 4, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to the LO flaperons. For example, a failure in the central CAN bus 1 would likely result in the LO flaperon going offline, leading to a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to the LM and RM rudder verso such that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the central CAN bus 1 would likely result in both the LM and RM rudder verso going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0045】 Similarly, referring to Figure 8C, the central CAN bus 2 may provide control signals to inverters 1 and 12, as well as 5 and 8, such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the central CAN bus 2 would result in both EPU 1 and EPU 12 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the central CAN bus 2 would result in both EPU 5 and EPU 8 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to TPAC 1 and TPAC 5. For example, a failure of the central CAN bus 2 would result in a loss of control of TPAC 1 and TPAC 5, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to the LO flaperons. For example, a failure in the central CAN bus 2 would likely result in the LO flaperon going offline, leading to a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to the LM and RM rudder verso such that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the central CAN bus 2 would likely result in both the LM and RM rudder verso going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0046】 Similarly, referring to Figure 8D, the right CAN bus 1 may provide control signals to inverters 3 and 10, as well as 5 and 8, such that a failure of this bus could result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the right CAN bus 1 would result in both EPU 3 and EPU 10 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the right CAN bus 1 would result in both EPU 5 and EPU 8 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to TPAC 3 and TPAC 5, such that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure of the right CAN bus 1 would result in a loss of control of TPAC 3 and TPAC 5, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to the RO flaperon. For example, a failure in the right CAN bus 1 would likely result in the RO flaperon going offline, leading to a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to the LI rudder and RI rudder, such that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the right CAN bus 1 would likely result in both the LI and RI rudder going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0047】 Similarly, referring to Figure 8D, the right CAN bus 2 may provide control signals to inverters 2 and 11, as well as 6 and 7, such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the right CAN bus 2 would result in both EPU 2 and EPU 11 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the right CAN bus 2 would result in both EPU 6 and EPU 7 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 2 may provide control signals to TPAC 2 and TPAC 6. For example, a failure of the right CAN bus 2 would result in a loss of control of TPAC 2 and TPAC 6, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 2 may provide control signals to the RO flaperon. For example, a failure in the right CAN bus 2 would likely result in the RO flaperon going offline, leading to a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 2 may provide control signals to the LI rudder and RI rudder, such that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the right CAN bus 2 would likely result in both the LI and RI rudder going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0048】 In some embodiments, the FCC may be configured to modify the FCC's control assignments to minimize, offset, or eliminate the loss of lift and / or control in the event of a failure of the left CAN bus 1. For example, referring to Figure 8B, if a failure of the left CAN bus 1 results in a loss of control of the BMS1, the FCC may modify the FCC's commands to the BMS4 via the left CAN bus 2 to minimize, offset, or eliminate the loss of lift in the event of a loss of control of the BMS1. As another example, if a failure of the left CAN bus 1 results in a loss of control of TPAC1 and TPAC6, the FCC may modify the FCC's commands in the event of a loss of control of TPAC1 and TPAC6 in relation to the loss of control of TPAC1 and TPAC6. Loss of lift Minimize it, Offset or Also is excluded To eliminate this, the FCC commands to TPAC3 and TPAC4 via left CAN bus 2 may be modified. Generally, the FCC, in the event of a failure in any of the CAN buses, is controlled by the control of any of the other CAN buses, and / or any control surface and / or controller associated with any of the other CAN buses. Loss of lift and / or control Minimize it, Offset or Also is excluded It may be configured to remove. In some embodiments, the FCC may be configured to modify any control margin associated with an inceptor input provided by the aircraft pilot. 【0049】 It should be understood that all such symmetrical arrangements and combinations of the EPU inverter and associated propeller blades, TPAC, BMS, flaperon and associated CSA, and control signaling to the rudder vater and associated CSA are contemplated within the scope of this disclosure. Furthermore, asymmetrical arrangements and combinations of the EPU inverter and associated propeller blades, TPAC, BMS, flaperon and associated CSA, and control signaling to the rudder vater and associated CSA are contemplated within the scope of this disclosure where necessary to provide safe or redundant control. For example, some asymmetrical arrangements may be required for a given total number of devices, each device having two connections, and symmetrical arrangements are preferred. 【0050】 With respect to the embodiment of Figure 8A, the configuration described above may offer the benefit that the aircraft remains controllable even in the event of (1) the loss of any FCC, or (2) the loss of all three CAN buses: left CAN bus 1, central CAN bus 1, and right CAN bus 1, or (3) the loss of all three CAN buses: left CAN bus 2, central CAN bus 2, and right CAN bus 2. Furthermore, minimally acceptable control may be achieved in such a configuration even after (1) the loss of any of the four flight control computer lanes: L FCC-A, L FCC-B, R FCC-A, or R FCC-B, or (2) the loss of any two of the CAN buses, or (3) the complete loss of the Aircraft Low Voltage Electrical Wiring Interconnect System (EWIS) channel and additional CAN buses. 【0051】 In some embodiments, minimally acceptable control may be defined as the condition under which the aircraft can continue flying for a minimum amount of time, e.g., 5, 10, 15, 20, 25, 30, 35, 40, or 45 minutes. In some embodiments, minimally acceptable control during thrustbone flight may include at least five controls from the six diagonal pairs of the EPU and associated TPAC. In some embodiments, minimally acceptable control during winged flight may include two of the three left-side tilter engines, two of the three right-side tilter engines, one from the two left-wing flaperons, one from the two right-wing flaperons, two from the three left-side rudder batons, and two from the three right-side rudder batons. 【0052】 Figure 9A illustrates another exemplary flight control signaling architecture for controlling a control plane and associated actuators in various embodiments. In the embodiments illustrated, L FCC-A may provide control signals via left CAN bus 1 and left CAN bus 2, and L FCC-B may provide control signals via central CAN bus 1 and central CAN bus 2 Control signals can be provided via the right CAN bus 1 and the right CAN bus 2. 【0053】 For example, referring to Figure 9B, the left CAN bus 1 may provide control signals to inverters 1, 3, 7, and 12 such that a failure of this bus may result in a substantially asymmetrical loss of lift or thrust on either side of the aircraft. For example, a failure of the left CAN bus 1 would result in EPUs 1, 3, 7, and 12 all going offline simultaneously, which would likely result in a substantially asymmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 1 may provide control signals to TPAC 1 and TPAC 3. For example, a failure of the left CAN bus 1 would result in a loss of control of TPAC 1 and TPAC 3, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's yaw and / or pitch. Furthermore, the left CAN bus 1 may provide control signals to the LO rudder and LM rudder. For example, a failure of the left CAN bus 1 would result in both the LO rudder and LM rudder going offline simultaneously, which would likely result in a substantially asymmetrical loss of lift with respect to the aircraft's yaw and / or pitch. 【0054】 For example, referring to Figure 9B, the left CAN bus 2 may provide control signals to inverters 4, 6, 9, and 10 such that a failure of this bus could result in a substantially asymmetrical loss of lift or thrust on either side of the aircraft. For example, a failure of the left CAN bus 2 would result in EPUs 4, 6, 9, and 10 all going offline simultaneously, which would likely result in a substantially asymmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 2 may provide control signals to TPACs 4 and 6. For example, a failure of the left CAN bus 2 would result in a loss of control of TPACs 4 and 6, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 2 may provide control signals to the LI flaperon and RI flaperon such that a failure of this bus could result in a substantially symmetrical loss of control on either side of the aircraft. For example, a failure of the left CAN bus 2 would result in both the LI flaperon and RI flaperon going offline simultaneously, which would likely result in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 2 may provide control signals to the RM rudder and RO rudder. For example, a failure in the left CAN bus 2 would likely result in both the RM and RO rudder going offline simultaneously, leading to a substantially asymmetrical loss of lift with respect to the aircraft's yaw and / or pitch. 【0055】 Furthermore, referring to Figure 9B, left CAN bus 1 may provide control signals to inverters 1, 3, 7, and 12, and left CAN bus 2 may provide control signals to inverters 4, 6, 9, and 10, so that a failure in either bus may result in a substantially symmetrical loss of lift or thrust on either side of the aircraft. For example, a failure in both left CAN bus 1 and left CAN bus 2 would result in EPUs 1 and 12, 3 and 10, 4 and 9, and 6 and 7 all going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, left CAN bus 1 may provide control signals to TPAC 1 and TPAC 3, and left CAN bus 2 may provide control signals to TPAC 4 and TPAC 6, so that a failure in either bus may result in a substantially symmetrical loss of control on either side of the aircraft. For example, a failure in both left CAN bus 1 and left CAN bus 2 would result in a loss of control of TPAC1 and 3, and TPAC4 and 6, respectively, likely resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, left CAN bus 1 may provide control signals to the LO rudder and LM rudder, and left CAN bus 2 may provide control signals to the RM rudder and RO rudder, so that a failure in either bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in both left CAN bus 1 and left CAN bus 2 would result in the LO rudder, LM rudder, RM rudder, and RO rudder all going offline simultaneously, likely resulting in a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0056】 For example, referring to Figure 9C, the central CAN bus 1 may provide control signals to inverters 1, 2, 9, and 11 such that a failure of this bus may result in a substantially asymmetrical loss of lift or thrust on either side of the aircraft. For example, a failure of the central CAN bus 1 would result in EPUs 1, 2, 9, and 11 all going offline simultaneously, which would likely result in a substantially asymmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to TPAC1 and TPAC2. For example, a failure of the central CAN bus 1 would result in a loss of control of TPAC1 and TPAC2, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to the LI flaperon and RO flaperon. For example, a failure of the central CAN bus 1 would result in both the LI flaperon and RO flaperon going offline simultaneously, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to the LM rudder and LI rudder. For example, a failure of the central CAN bus 1 would likely result in both the LM and LI rudder going offline simultaneously, leading to a substantially asymmetrical loss of lift with respect to the aircraft's yaw and / or pitch. 【0057】 For example, referring to Figure 9C, the central CAN bus 2 may provide control signals to inverters 4, 5, 8, and 12 such that a failure of this bus may result in a substantially asymmetrical loss of lift or thrust on either side of the aircraft. For example, a failure of the central CAN bus 2 would result in EPUs 4, 5, 8, and 12 all going offline simultaneously, which would likely result in a substantially asymmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to TPACs 4 and 5. For example, a failure of the central CAN bus 2 would result in a loss of control of TPACs 4 and 5, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to the LO flaperons. For example, a failure of the central CAN bus 2 would result in the LO flaperons going offline simultaneously, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to the RI rudder and RM rudder. For example, a failure in the central CAN bus 2 would likely result in both the RI rudder and RM rudder going offline simultaneously, leading to a substantially asymmetrical loss of lift with respect to the aircraft's yaw and / or pitch. 【0058】 Furthermore, referring to Figure 9C, central CAN bus 1 may provide control signals to inverters 1, 2, 9, and 12, and central CAN bus 2 may provide control signals to inverters 4, 5, 8, and 12, so that a failure in either bus may result in a substantially symmetrical loss of lift or thrust on either side of the aircraft. For example, a failure in both central CAN bus 1 and central CAN bus 2 would result in EPUs 1 and 12, 2 and 11, 4 and 9, and 5 and 8 all going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, central CAN bus 1 may provide control signals to TPAC 1 and TPAC 2, and central CAN bus 2 may provide control signals to TPAC 4 and TPAC 5. For example, a failure in both central CAN bus 1 and central CAN bus 2 would result in a loss of control of TPAC 1 and 2, and TPAC 4 and 5, respectively, which is likely to result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, central CAN bus 1 may provide control signals to the LI and RO flaperons, and central CAN bus 2 may provide control signals to the LO flaperon. For example, a failure in both central CAN bus 1 and central CAN bus 2 would result in the LO, LI, and RO flaperons all going offline simultaneously, likely resulting in a substantially asymmetric loss of control with respect to the aircraft's roll and / or pitch. Additionally, central CAN bus 1 may provide control signals to the LM and LI ruddervers, and central CAN bus 2 may provide control signals to the RI and RM ruddervers, so that a failure in either bus could result in a substantially symmetric loss of control on both sides of the aircraft. For example, a failure in both central CAN bus 1 and central CAN bus 2 would result in the LM, LI, RI, and RM ruddervers all going offline simultaneously, likely resulting in a substantially symmetric loss of control with respect to the aircraft's yaw and / or pitch. 【0059】 For example, referring to Figure 9D, the right CAN bus 1 may provide control signals to inverters 2, 3, 8, and 10 such that a failure of this bus may result in a substantially asymmetrical loss of lift or thrust on either side of the aircraft. For example, a failure of the right CAN bus 1 would result in EPUs 2, 3, 8, and 10 all going offline simultaneously, which would likely result in a substantially asymmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to TPACs 2 and 3. For example, a failure of the right CAN bus 1 would result in a loss of control of TPACs 2 and 3, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to the LO flaperon and RI flaperon. For example, a failure of the right CAN bus 1 would result in both the LO flaperon and RI flaperon going offline simultaneously, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, right CAN bus 1 may provide control signals to the LO rudder and LI rudder. For example, a failure of right CAN bus 1 would likely result in both the LO and LI rudder going offline simultaneously, leading to a substantially asymmetrical loss of lift with respect to the aircraft's yaw and / or pitch. 【0060】 For example, referring to Figure 9D, the right CAN bus 2 may provide control signals to inverters 5, 6, 7, and 11 such that a failure of this bus may result in a substantially asymmetrical loss of lift or thrust on either side of the aircraft. For example, a failure of the central CAN bus 2 would result in EPUs 5, 6, 7, and 11 all going offline simultaneously, which is likely to result in a substantially asymmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to TPACs 5 and TPACs 6. For example, a failure of the right CAN bus 2 would result in a loss of control of TPACs 5 and TPACs 6, which is likely to result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 2 may provide control signals to the RO flaperon. For example, a failure of the right CAN bus 2 would result in the RO flaperon ga O This would result in the aircraft flying, likely leading to a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, right CAN bus 2 may provide control signals to the RI rudder and RO rudder. For example, a failure of right CAN bus 2 would likely result in both the RI rudder and RO rudder going offline simultaneously, likely leading to a substantially asymmetrical loss of lift with respect to the aircraft's yaw and / or pitch. 【0061】 Furthermore, referring to Figure 9D, right CAN bus 1 may provide control signals to inverters 2, 3, 8, and 10, and right CAN bus 2 may provide control signals to inverters 5, 6, 7, and 11, so that a failure in either bus may result in a substantially symmetrical loss of lift or thrust on either side of the aircraft. For example, a failure in both right CAN bus 1 and right CAN bus 2 would result in EPUs 2, 11, 3, 10, 5, 8, 6, and 7 all going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. In addition, right CAN bus 1 may provide control signals to TPACs 2 and 3, and right CAN bus 2 may provide control signals to TPACs 5 and 6. For example, a failure in both right CAN bus 1 and right CAN bus 2 would result in a loss of control of TPACs 2 and 3, and TPACs 5 and 6, respectively, which is likely to result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, right CAN bus 1 may provide control signals to the LO and RI flaperons, and right CAN bus 2 may provide control signals to the RO flaperon. For example, a failure in both right CAN bus 1 and right CAN bus 2 would result in the LO, RI, and RO flaperons all going offline simultaneously, likely resulting in a substantially asymmetric loss of control with respect to the aircraft's roll and / or pitch. Furthermore, right CAN bus 1 may provide control signals to the LO and LI ruddervers, and right CAN bus 2 may provide control signals to the RI and RO ruddervers, so that a failure in both buses could result in a substantially symmetric loss of control on both sides of the aircraft. For example, a failure in both right CAN bus 1 and right CAN bus 2 would result in the LO, LI, RI, and RO ruddervers all going offline simultaneously, likely resulting in a substantially symmetric loss of control with respect to the aircraft's yaw and / or pitch. 【0062】 With respect to the embodiment of Figure 9A, such a configuration may offer the benefit that the aircraft remains fully controllable even in the event of (1) any loss of the FCC or (2) any loss of the CAN bus. Furthermore, minimally acceptable control may be achieved in such a configuration even after the loss of any two of the CAN buses. 【0063】 With respect to Figure 9A, in the alternative embodiment, L FCC-A provides control signals via left CAN bus 1 and left CAN bus 2, and L FCC-B provides control signals via central CAN bus 1 and central CAN bus 2 An alternative configuration may exist in which, instead of providing control signals via the FCC-A and R FCC-A providing control signals via the right CAN bus 1 and right CAN bus 2, L FCC-A providing control signals via the left CAN bus 1 and left CAN bus 2, L FCC-B providing control signals via the right CAN bus 1, R FCC-A providing control signals via the central CAN bus 1 and central CAN bus 2, and R FCC-B providing control signals via the right CAN bus 2. This proposed alternative embodiment may offer the benefit that the aircraft remains fully controllable even with (1) the loss of any FCC or (2) the loss of any CAN bus. Furthermore, minimally acceptable control may be achieved in such a configuration even after (1) the loss of any two of the CAN buses. 【0064】 Furthermore, with respect to Figure 9A, there may be yet another alternative embodiment in which L FCC-A may provide control signals via left CAN bus 1 and central CAN bus 2, L FCC-B may provide control signals via left CAN bus 2 and right CAN bus 2, R FCC-A may provide control signals via central CAN bus 1 and right CAN bus 2, and R FCC-B may provide control signals via left CAN bus 1 and right CAN bus 1. This proposed alternative embodiment may offer the benefit that (1) the aircraft remains fully controllable even with the loss of any CAN bus. Furthermore, minimally acceptable control may be achieved in such a configuration even after (1) the loss of flight control computer lanes L FCC-A and L FCC-B, (2) the loss of flight control computer lanes R FCC-A and R FCC-B, (3) the loss of flight control computer lanes L FCC-A and R FCC-B, (4) the loss of flight control computer lanes L FCC-B and R FCC-B, or (5) the loss of any two of the CAN buses. 【0065】 Figure 10A illustrates another exemplary flight control signaling architecture for controlling a control surface and associated actuators according to various embodiments. In the embodiments illustrated, L FCC-A may provide control signals via left CAN bus 1 and center CAN bus 2, L FCC-B may provide control signals via left CAN bus 2 and right CAN bus 2, R FCC-A may provide control signals via center CAN bus 1 and right CAN bus 2, and R FCC-B may provide control signals via left CAN bus 1 and right CAN bus 1. 【0066】 For example, referring to Figure 10B, the left CAN bus 1 may provide control signals to inverters 1, 3, 7, and 12 such that a failure of this bus could result in a substantially asymmetrical loss of lift or thrust on either side of the aircraft. For example, a failure of the left CAN bus 1 would result in EPUs 1, 3, 7, and 12 all going offline simultaneously, which would likely result in a substantially asymmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 1 may provide control signals to TPAC1 and TPAC3. For example, a failure of the left CAN bus 1 would result in a loss of control of TPAC1 and TPAC3, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 1 may provide control signals to the LI flaperon and RI flaperon such that a failure of this bus could result in a substantially symmetrical loss of control on either side of the aircraft. For example, a failure of the left CAN bus 1 would result in both the LI flaperon and RI flaperon going offline simultaneously, which would likely result in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 1 may provide control signals to the LM and RM rudder vortices in such a way that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure of the left CAN bus 1 would likely cause both the LM and RM rudder vortices to go offline simultaneously, resulting in a substantially symmetrical loss of lift with respect to the aircraft's yaw and / or pitch. 【0067】 For example, referring to Figure 10B, the left CAN bus 2 may provide control signals to inverters 4, 6, 9, and 10 such that a failure of this bus could result in a substantially asymmetrical loss of lift or thrust on either side of the aircraft. For example, a failure of the left CAN bus 2 would result in EPUs 4, 6, 9, and 10 all going offline simultaneously, which would likely result in a substantially asymmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 2 may provide control signals to TPACs 4 and 6. For example, a failure of the left CAN bus 2 would result in a loss of control of TPACs 4 and 6, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 2 may provide control signals to the LI flaperon and RI flaperon such that a failure of this bus could result in a substantially symmetrical loss of control on either side of the aircraft. For example, a failure of the left CAN bus 2 would result in both the LI flaperon and RI flaperon going offline simultaneously, which would likely result in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 2 may provide control signals to the LM and RM rudder vortices in such a way that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure of the left CAN bus 2 would likely cause both the LM and RM rudder vortices to go offline simultaneously, resulting in a substantially symmetrical loss of lift with respect to the aircraft's yaw and / or pitch. 【0068】 Furthermore, referring to Figure 10B, left CAN bus 1 may provide control signals to inverters 1, 3, 7, and 12, and left CAN bus 2 may provide control signals to inverters 4, 6, 9, and 10, so that a failure in either bus may result in a substantially symmetrical loss of lift or thrust on either side of the aircraft. For example, a failure in both left CAN bus 1 and left CAN bus 2 would result in EPUs 1 and 12, 3 and 10, 4 and 9, and 6 and 7 all going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, left CAN bus 1 may provide control signals to TPAC 1 and TPAC 3, and left CAN bus 2 may provide control signals to TPAC 4 and TPAC 6, so that a failure in either bus may result in a substantially symmetrical loss of control on either side of the aircraft. For example, a failure in both left CAN bus 1 and left CAN bus 2 would result in a loss of control of TPAC1 and 3, and TPAC4 and 6, respectively, likely resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, both left CAN bus 1 and left CAN bus 2 may provide control signals to the LI flaperon and RI flaperon so that a failure in either bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in both left CAN bus 1 and left CAN bus 2 would result in both the LI flaperon and RI flaperon going offline simultaneously, likely resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, both left CAN bus 1 and left CAN bus 2 may provide control signals to the LM rudder and RM rudder, so that a failure in either bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in both left CAN bus 1 and left CAN bus 2 would likely result in both the LM rudder and RM rudder going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0069】 For example, referring to Figure 10C, the central CAN bus 1 may provide control signals to inverters 1, 2, 9, and 11 such that a failure of this bus may result in a substantially asymmetrical loss of lift or thrust on either side of the aircraft. For example, a failure of the central CAN bus 1 would result in EPUs 1, 2, 9, and 11 all going offline simultaneously, which is likely to result in a substantially asymmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to TPAC1 and TPAC2. For example, a failure of the central CAN bus 1 would result in a loss of control of TPAC1 and TPAC2, which is likely to result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to the RO flaperon. For example, a failure of the central CAN bus 1 would result in the RO flaperon ga O This would result in the aircraft flying, which is likely to cause a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to the LI rudder and RI rudder. For example, a failure in the central CAN bus 1 would likely cause both the LI rudder and RI rudder to go offline simultaneously, resulting in a substantially symmetrical loss of lift with respect to the aircraft's yaw and / or pitch. 【0070】 For example, referring to Figure 10C, the central CAN bus 2 may provide control signals to inverters 4, 5, 8, and 12 such that a failure of this bus may result in a substantially asymmetrical loss of lift or thrust on either side of the aircraft. For example, a failure of the central CAN bus 2 would result in all EPUs 4, 5, 8, and 12 going offline simultaneously, which would likely result in a substantially asymmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to TPACs 4 and 5. For example, a failure of the central CAN bus 2 would result in a loss of control of TPACs 4 and 5, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to the RO flaperon. For example, a failure of the central CAN bus 2 would result in the RO flaperon ga O This would result in the aircraft flying, which is likely to cause a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to the LI rudder and RI rudder, such that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure of the central CAN bus 2 is likely to cause both the LI rudder and RI rudder to go offline simultaneously, resulting in a substantially symmetrical loss of lift with respect to the aircraft's yaw and / or pitch. 【0071】 Furthermore, referring to Figure 10C, central CAN bus 1 may provide control signals to inverters 1, 2, 9, and 12, and central CAN bus 2 may provide control signals to inverters 4, 5, 8, and 12, so that a failure in either bus may result in a substantially symmetrical loss of lift or thrust on either side of the aircraft. For example, a failure in both central CAN bus 1 and central CAN bus 2 would result in EPUs 1 and 12, 2 and 11, 4 and 9, and 5 and 8 all going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, central CAN bus 1 may provide control signals to TPAC 1 and TPAC 2, and central CAN bus 2 may provide control signals to TPAC 4 and TPAC 5. For example, a failure in both central CAN bus 1 and central CAN bus 2 would result in a loss of control of TPAC 1 and 2, and TPAC 4 and 5, respectively, which is likely to result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, both Central CAN Bus 1 and Central CAN Bus 2 can provide control signals to the RO flaperon. For example, a failure in both Central CAN Bus 1 and Central CAN Bus 2 will cause the RO flaperon to fail. ga O This would result in the aircraft flying, which is likely to result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, both central CAN bus 1 and central CAN bus 2 can provide control signals to the LI rudder and RI rudder, such that a failure in either bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in both central CAN bus 1 and central CAN bus 2 would result in both the LI rudder and RI rudder going offline simultaneously, which is likely to result in a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0072】 For example, referring to Figure 10D, the right CAN bus 1 may provide control signals to inverters 2, 3, 8, and 10 such that a failure of this bus may result in a substantially asymmetrical loss of lift or thrust on either side of the aircraft. For example, a failure of the right CAN bus 1 would result in EPUs 2, 3, 8, and 10 all going offline simultaneously, which would likely result in a substantially asymmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to TPACs 2 and 3. For example, a failure of the right CAN bus 1 would result in a loss of control of TPACs 2 and 3, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to the LO flaperons. For example, a failure of the right CAN bus 1 would result in the LO flaperons going offline simultaneously, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to the LO rudder and RO rudder, such that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure of the right CAN bus 1 would result in both the LO rudder and RO rudder going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's yaw and / or pitch. 【0073】 For example, referring to Figure 10D, the right CAN bus 2 may provide control signals to inverters 5, 6, 7, and 11 such that a failure of this bus may result in a substantially asymmetrical loss of lift or thrust on either side of the aircraft. For example, a failure of the central CAN bus 2 would result in EPUs 5, 6, 7, and 11 all going offline simultaneously, which would likely result in a substantially asymmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to TPACs 5 and 6. For example, a failure of the right CAN bus 2 would result in a loss of control of TPACs 5 and 6, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 2 may provide control signals to the LO flaperons. For example, a failure of the right CAN bus 2 would result in the LO flaperons going offline simultaneously, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 2 may provide control signals to the LO rudder and RO rudder, such that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure of the right CAN bus 2 would result in both the LO rudder and RO rudder going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's yaw and / or pitch. 【0074】 Furthermore, referring to Figure 10D, right CAN bus 1 may provide control signals to inverters 2, 3, 8, and 10, and right CAN bus 2 may provide control signals to inverters 5, 6, 7, and 11, so that a failure in either bus may result in a substantially symmetrical loss of lift or thrust on either side of the aircraft. For example, a failure in both right CAN bus 1 and right CAN bus 2 would result in EPUs 2, 11, 3, 10, 5, 8, 6, and 7 all going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, right CAN bus 1 may provide control signals to TPACs 2 and 3, and right CAN bus 2 may provide control signals to TPACs 5 and 6. For example, a failure in both right CAN bus 1 and right CAN bus 2 would result in a loss of control of TPACs 2 and 3, and TPACs 5 and 6, respectively, which is likely to result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, both right CAN bus 1 and right CAN bus 2 can provide control signals to the RO flaperon. For example, a failure in both right CAN bus 1 and right CAN bus 2 will cause the RO flaperon to fail. ga O This would result in the aircraft flying offline, which is likely to result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, both right CAN bus 1 and right CAN bus 2 can provide control signals to the LO rudder and RO rudder, such that a failure in either bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in both right CAN bus 1 and right CAN bus 2 would result in both the LO rudder and RO rudder going offline simultaneously, which is likely to result in a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0075】 With respect to the embodiment shown in Figure 10A, the configuration described above may offer the benefit that minimally acceptable control of the aircraft can be achieved in such a configuration even after (1) the loss of any two CAN buses, (2) the loss of flight control computer lanes L FCC-A and L FCC-B, or (3) the loss of flight control computer lanes R FCC-A and R FCC-B. 【0076】 Figure 11A illustrates another exemplary flight control signaling architecture for controlling a control surface and associated actuators according to various embodiments. In the embodiments illustrated, L FCC-A may provide control signals via left CAN bus 1 and center CAN bus 2, L FCC-B may provide control signals via left CAN bus 2 and right CAN bus 1, R FCC-A may provide control signals via center CAN bus 1 and right CAN bus 2, and R FCC-B may provide control signals via left CAN bus 1 and right CAN bus 2. 【0077】 For example, referring to Figure 11B, the left CAN bus 1 may provide control signals to inverters 1 and 12, as well as 3 and 10, such that a failure of this bus could result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the left CAN bus 1 would result in both EPU 1 and EPU 12 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the left CAN bus 1 would result in both EPU 3 and EPU 10 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 1 may provide control signals to TPAC 1 and TPAC 3. For example, a failure of the left CAN bus 1 would result in a loss of control of TPAC 1 and TPAC 3, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 1 may provide control signals to the LI flaperon and RI flaperon, such that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in left CAN bus 1 is likely to cause both the LI and RI flaperons to go offline simultaneously, resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, left CAN bus 1 may provide control signals to the LM and RM rudder verso such that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in left CAN bus 1 is likely to cause both the LM and RM rudder verso to go offline simultaneously, resulting in a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0078】 Similarly, referring to Figure 11B, the left CAN bus 2 may provide control signals to inverters 4 and 9, as well as 6 and 7, such that a failure of this bus could result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the left CAN bus 2 would result in both EPU 4 and EPU 9 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the left CAN bus 2 would result in both EPU 6 and EPU 7 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 2 may provide control signals to TPAC 4 and TPAC 6. For example, a failure of the left CAN bus 2 would result in a loss of control of TPAC 4 and TPAC 6, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 2 may provide control signals to the LI flaperon and RI flaperon, such that a failure of this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the left CAN bus 2 is likely to cause both the LI and RI flaperons to go offline simultaneously, resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 2 may provide control signals to the LM and RM rudder verso such that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the left CAN bus 2 is likely to cause both the LM and RM rudder verso to go offline simultaneously, resulting in a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0079】 Furthermore, referring to Figure 11B, left CAN bus 1 may provide control signals to inverters 1, 3, 10, and 12, and left CAN bus 2 may provide control signals to inverters 4, 6, 7, and 9, so that a failure in either bus may result in a substantially symmetrical loss of lift or thrust on either side of the aircraft. For example, a failure in both left CAN bus 1 and left CAN bus 2 would result in EPUs 1 and 12, 3 and 10, 4 and 7, and 6 and 7 all going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. In addition, left CAN bus 1 may provide control signals to TPAC 1 and TPAC 3, and left CAN bus 2 may provide control signals to TPAC 4 and TPAC 6, so that a failure in either bus may result in a substantially symmetrical loss of control on either side of the aircraft. For example, a failure in both left CAN bus 1 and left CAN bus 2 would result in a loss of control of TPAC1 and 3, and TPAC4 and 6, respectively, likely resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, both left CAN bus 1 and left CAN bus 2 can provide control signals to the LI flaperon and RI flaperon so that a failure in either bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in both left CAN bus 1 and left CAN bus 2 would result in all LI and RI flaperon going offline simultaneously, likely resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, both left CAN bus 1 and left CAN bus 2 can provide control signals to the LM rudder and RM rudder, so that a failure in either bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in both left CAN bus 1 and left CAN bus 2 would likely result in both the LM rudder and RM rudder going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0080】 Similarly, referring to Figure 11C, the central CAN bus 1 may provide control signals to inverters 1 and 12, as well as 2 and 11, such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the central CAN bus 1 would result in both EPU 1 and EPU 12 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the central CAN bus 1 would result in both EPU 2 and EPU 11 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to TPAC 1 and TPAC 2. For example, a failure of the central CAN bus 1 would result in a loss of control of TPAC 1 and TPAC 2, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to the RO flaperon. For example, a failure in the central CAN bus 1 would likely result in the RO flaperon going offline, leading to a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to the LI rudder and RI rudder, such that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the central CAN bus 1 would likely result in both the LI and RI rudder going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0081】 Similarly, referring to Figure 11C, the central CAN bus 2 may provide control signals to inverters 4 and 9, as well as 5 and 8, such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the central CAN bus 2 would result in both EPU 4 and EPU 9 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the central CAN bus 2 would result in both EPU 5 and EPU 8 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to TPAC 4 and TPAC 5. For example, a failure of the central CAN bus 2 would result in a loss of control of TPAC 4 and TPAC 5, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to the RO flaperon. For example, a failure in the central CAN bus 2 would likely result in the RO flaperon going offline, leading to a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to the LI rudder and RI rudder, such that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the central CAN bus 2 would likely result in both the LI and RI rudder going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0082】 Furthermore, referring to Figure 11C, central CAN bus 1 may provide control signals to inverters 1, 2, 11, and 12, and central CAN bus 2 may provide control signals to inverters 4, 5, 8, and 9, so that a failure in either bus may result in a substantially symmetrical loss of lift or thrust on either side of the aircraft. For example, a failure in both central CAN bus 1 and central CAN bus 2 would result in EPUs 1 and 12, 2 and 11, 4 and 9, and 5 and 8 all going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, central CAN bus 1 may provide control signals to TPAC 1 and TPAC 2, and central CAN bus 2 may provide control signals to TPAC 4 and TPAC 5. For example, a failure in central CAN bus 1 and central CAN bus 2 would result in a loss of control of TPAC 1 and 2, and TPAC 4 and 5, respectively, which is likely to result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, both Central CAN Bus 1 and Central CAN Bus 2 can provide control signals to the RO flaperon. For example, a failure in both Central CAN Bus 1 and Central CAN Bus 2 will cause the RO flaperon to fail. ga O This would result in the aircraft flying, which is likely to result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, both central CAN bus 1 and central CAN bus 2 can provide control signals to the LI rudder and RI rudder, such that a failure in either bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in both central CAN bus 1 and central CAN bus 2 would result in both the LI rudder and RI rudder going offline simultaneously, which is likely to result in a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0083】 Similarly, referring to Figure 11D, the right CAN bus 1 may provide control signals to inverters 2 and 11, as well as 3 and 10, such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the right CAN bus 1 would result in both EPU 2 and EPU 11 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the right CAN bus 1 would result in both EPU 3 and EPU 10 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to TPAC 2 and TPAC 3. For example, a failure of the right CAN bus 1 would result in a loss of control of TPAC 2 and TPAC 3, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to the LO flaperon. For example, a failure in the right CAN bus 1 would likely result in the LO flaperon going offline, leading to a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to the LO rudder and RO rudder, such that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the right CAN bus 1 would likely result in both the LO rudder and RO rudder going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0084】 Similarly, referring to Figure 11D, the right CAN bus 2 may provide control signals to inverters 5 and 8, as well as 6 and 7, such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the right CAN bus 1 would result in both EPU 5 and EPU 8 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the right CAN bus 2 would result in both EPU 6 and EPU 7 going offline simultaneously, which would likely result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 2 may provide control signals to TPAC 5 and TPAC 6. For example, a failure of the right CAN bus 2 would result in a loss of control of TPAC 5 and TPAC 6, which would likely result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 2 may provide control signals to the LO flaperon. For example, a failure in the right CAN bus 2 would likely result in the LO flaperon going offline, leading to a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 2 may provide control signals to the LO rudder and RO rudder, such that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the right CAN bus 2 would likely result in both the LO rudder and RO rudder going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0085】 Furthermore, referring to Figure 11D, right CAN bus 1 may provide control signals to inverters 2, 3, 10, and 11, and right CAN bus 2 may provide control signals to inverters 5, 6, 7, and 8, so that a failure in either bus may result in a substantially symmetrical loss of lift or thrust on either side of the aircraft. For example, a failure in both right CAN bus 1 and right CAN bus 2 would result in EPUs 2 and 11, 3 and 10, 5 and 8, and 6 and 7 all going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, right CAN bus 1 may provide control signals to TPAC 2 and TPAC 3, and right CAN bus 2 may provide control signals to TPAC 5 and TPAC 6. For example, a failure in right CAN bus 1 and right CAN bus 2 would result in a loss of control of TPAC 2 and 3, and TPAC 5 and 6, respectively, which is likely to result in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, both right CAN bus 1 and right CAN bus 2 can provide control signals to the LO flaperon. For example, a failure in both right CAN bus 1 and right CAN bus 2 would likely result in the LO flaperon going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Additionally, both right CAN bus 1 and right CAN bus 2 can provide control signals to the LO rudder and RO rudder, such that a failure in either bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in both right CAN bus 1 and right CAN bus 2 would likely result in the LO rudder and RO rudder all going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0086】 With respect to the embodiment of Figure 11A, the configuration described above may offer the benefit that minimally acceptable control of the aircraft can be achieved even after (1) loss of any two CAN buses, (2) loss of any two FCCs, or (3) loss of any two Aircraft Low Voltage Electrical Wiring Interconnection System (EWIS) channels has no substantial impact on the inverter. 【0087】 Figure 12A illustrates another exemplary flight control signaling architecture for controlling a control plane and associated actuators according to various embodiments. In the embodiments illustrated, L FCC-A may provide control signals via left CAN bus 1, L FCC-B may provide control signals via right CAN bus 1, R FCC-A may provide control signals via center CAN bus 1, and R FCC-B may provide control signals via right CAN bus 2. 【0088】 For example, referring to Figure 12B, the left CAN bus 1 may provide control signals to inverters 1 and 12, 4 and 9, and 6 and 7, such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the left CAN bus 1 would result in both EPU 1 and EPU 12 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the left CAN bus 1 would result in both EPU 4 and EPU 9 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the left CAN bus 1 would result in both EPU 6 and EPU 7 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 1 may provide control signals to TPAC 1, TPAC 4, and TPAC 5. For example, a failure in left CAN bus 1 would result in a loss of control of TPACs 1, 4, and 6, likely resulting in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, left CAN bus 1 may provide control signals to the LI and RI flaperons so that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in left CAN bus 1 would likely cause both the LI and RI flaperons to go offline simultaneously, likely resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, left CAN bus 1 may provide control signals to the LM, RM, and RO ruddervers. For example, a failure in left CAN bus 1 would likely cause the LM, RM, and RO ruddervers to all go offline simultaneously, likely resulting in a substantially asymmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0089】 Similarly, referring to Figure 12C, the central CAN bus 1 may provide control signals to inverters 1 and 12, 2 and 11, and 5 and 8 such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the central CAN bus 1 would result in both EPU 1 and EPU 12 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the central CAN bus 1 would result in both EPU 2 and EPU 11 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the central CAN bus 1 would result in both EPU 5 and EPU 8 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 、T PAC1, TPAC2, and TPAC5 Control signals It may provide the following: For example, a failure of the central CAN bus 1 would result in a loss of control of TPACs 1, 2, and 5, likely resulting in a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to the LO and RO flaperons so that a failure of this bus may result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure of the central CAN bus 1 would likely cause both the LO and RO flaperons to go offline simultaneously, likely resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 1 may provide control signals to the LI, RI, and RM ruddervers. For example, a failure of the central CAN bus 1 would likely cause the LI, RI, and RM ruddervers to all go offline simultaneously, likely resulting in a substantially asymmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0090】 Similarly, referring to Figure 12D, the right CAN bus 1 may provide control signals to inverters 2 and 11, 3 and 10, and 4 and 9, such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the right CAN bus 1 would result in both EPU2 and EPU11 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the right CAN bus 1 would result in both EPU3 and EPU10 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the right CAN bus 1 would result in both EPU4 and EPU9 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to TPAC2, TPAC3, and TPAC4. For example, a failure in the right CAN bus 1 would likely result in a loss of control of TPACs 2, 3, and 4, leading to a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to the LI and RI flaperons, both of which would simultaneously go offline, so that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the right CAN bus 1 would likely cause both the LI and RI flaperons to simultaneously go offline, resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 may provide control signals to the LO, LM, and LI ruddervers. For example, a failure in the right CAN bus 1 would likely cause the LO, LM, and LI ruddervers to all simultaneously go offline, resulting in a substantially asymmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0091】 Similarly, referring to Figure 12D, the right CAN bus 2 may provide control signals to inverters 3 and 10, 5 and 8, and 6 and 7, such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the right CAN bus 2 would result in both EPU 3 and EPU 10 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the right CAN bus 2 would result in both EPU 5 and EPU 8 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the right CAN bus 2 would result in both EPU 6 and EPU 7 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 2 may provide control signals to TPAC 3, TPAC 5, and TPAC 6. For example, a failure in the right CAN bus 2 would likely result in a loss of control of TPACs 3, 5, and 6, leading to a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 2 may provide control signals to the LO and RO flaperons so that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the right CAN bus 2 would likely cause both the LO and RO flaperons to go offline simultaneously, resulting in a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 2 may provide control signals to the LO, RI, and RO ruddervers. For example, a failure in the right CAN bus 2 would likely cause the LO, RI, and RO ruddervers to all go offline simultaneously, resulting in a substantially asymmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0092】 With respect to the embodiment of Figure 12A, the configuration described above may offer the benefit that minimally acceptable control of the aircraft can be achieved even with (1) the loss of any two CAN buses or (2) the loss of any two FCCs. Furthermore, the embodiment of Figure 12A described above may offer the benefit of smaller wire weight compared to the disclosed embodiment utilizing six CAN buses. 【0093】 With respect to Figure 12A, in an alternative embodiment, instead of L FCC-A providing control signals via left CAN bus 1, L FCC-B providing control signals via right CAN bus 1, R FCC-A providing control signals via central CAN bus 1, and R FCC-B providing control signals via right CAN bus 2, there may be an alternative configuration in which L FCC-A provides control signals via left CAN bus 1 and central CAN bus 1, L FCC-B provides control signals via right CAN bus 1 and right CAN bus 2, R FCC-A provides control signals via central CAN bus 1 and right CAN bus 1, and R FCC-B provides control signals via left CAN bus 1 and right CAN bus 2. This proposed alternative embodiment may offer the benefit that the aircraft remains fully controllable even with (1) loss of any FCC or (2) loss of any CAN bus. Furthermore, minimally acceptable control may be achieved in such a configuration even after (1) loss of any two of the CAN buses. 【0094】 Figure 13A illustrates another exemplary flight control signaling architecture for controlling a control plane and associated actuators according to various embodiments. In the embodiments illustrated, L FCC-A may provide control signals via left CAN bus 1, L FCC-B may provide control signals via center CAN bus 2, R FCC-A may provide control signals via center CAN bus 1, and R FCC-B may provide control signals via right CAN bus 1. 【0095】 For example, referring to Figure 13B, the left CAN bus 1 may provide control signals to inverters 1 and 12, 4 and 9, and 6 and 7, such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the left CAN bus 1 would result in both EPU 1 and EPU 12 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the left CAN bus 1 would result in both EPU 4 and EPU 9 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the left CAN bus 1 would result in both EPU 6 and EPU 7 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the left CAN bus 1 may provide control signals to TPAC 1, TPAC 4, and TPAC 6. For example, a failure in left CAN bus 1 could result in TPAC 1, 4, and 6 all going offline simultaneously, potentially leading to a substantially asymmetric loss of control with respect to the aircraft's roll and / or pitch. Furthermore, left CAN bus 1 may provide control signals to the LI and RO flaperons. For example, a failure in left CAN bus 1 is likely to result in both the LI and RO flaperons going offline simultaneously, potentially leading to a substantially asymmetric loss of control with respect to the aircraft's roll and / or pitch. Furthermore, left CAN bus 1 may provide control signals to the LI, RI, and RO ruddervers. For example, a failure in left CAN bus 1 is likely to result in the LI, RI, and RO ruddervers going offline simultaneously, potentially leading to a substantially asymmetric loss of control with respect to the aircraft's yaw and / or pitch. 【0096】 Similarly, referring to Figure 13C, the central CAN bus 1 may provide control signals to inverters 1 and 12, 2 and 11, and 5 and 8 such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the central CAN bus 1 would result in both EPU 1 and EPU 12 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the central CAN bus 1 would result in both EPU 2 and EPU 11 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the central CAN bus 1 would result in both EPU 5 and EPU 8 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to TPAC 1, TPAC 2, and TPAC 5. For example, a failure in the central CAN bus 1 could result in TPAC 1, 2, and 5 all going offline simultaneously, leading to a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to the LO and RO flaperons so that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the central CAN bus 1 is likely to result in both the LO and RO flaperons going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 1 may provide control signals to the LO, LM, and RI ruddervers. For example, a failure in the central CAN bus 1 is likely to result in the LO, LM, and RI ruddervers all going offline simultaneously, leading to a substantially asymmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0097】 Similarly, referring to Figure 13C, the central CAN bus 2 may provide control signals to inverters 2 and 11, 3 and 10, and 4 and 9, such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the central CAN bus 2 would result in both EPU2 and EPU11 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the central CAN bus 2 would result in both EPU3 and EPU10 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the central CAN bus 2 would result in both EPU4 and EPU9 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to TPAC2, TPAC3, and TPAC4. For example, a failure in the central CAN bus 2 could result in TPAC2, 3, and 4 all going offline simultaneously, leading to a substantially asymmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to the LI and RI flaperons so that a failure in this bus could result in a substantially symmetrical loss of control on both sides of the aircraft. For example, a failure in the central CAN bus 2 is likely to result in both the LI and RI flaperons going offline simultaneously, leading to a substantially symmetrical loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the central CAN bus 2 may provide control signals to the LM, LI, and RM ruddervers. For example, a failure in the central CAN bus 2 is likely to result in the LM, LI, and RM ruddervers all going offline simultaneously, leading to a substantially asymmetrical loss of control with respect to the aircraft's yaw and / or pitch. 【0098】 Similarly, referring to Figure 13D, the right CAN bus 1 may provide control signals to inverters 3 and 10, 5 and 8, and 6 and 7, such that a failure of this bus may result in a substantially symmetrical loss of lift or thrust on both sides of the aircraft. For example, a failure of the right CAN bus 1 would result in both EPU 3 and EPU 10 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the right CAN bus 1 would result in both EPU 5 and EPU 8 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Similarly, a failure of the right CAN bus 1 would result in both EPU 6 and EPU 7 going offline simultaneously, which is likely to result in a substantially symmetrical loss of lift with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 、T PAC3, TPAC5, and TPAC6 Control signals It can provide control signals to the following: For example, a failure in the right CAN bus 1 could result in TPAC 3, 5, and 6 all going offline simultaneously, potentially leading to a substantially asymmetric loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 can provide control signals to the LO flaperon and RI flaperon. For example, a failure in the right CAN bus 1 could result in both the LO flaperon and RI flaperon going offline simultaneously, likely leading to a substantially asymmetric loss of control with respect to the aircraft's roll and / or pitch. Furthermore, the right CAN bus 1 can provide control signals to the LO rudder, RM rudder, and RO rudder. For example, a failure in the right CAN bus 1 could result in the LO rudder, RM rudder, and RO rudder all going offline simultaneously, likely leading to a substantially asymmetric loss of control with respect to the aircraft's yaw and / or pitch. 【0099】 With respect to the embodiment of Figure 13A, the configuration described above may offer the benefit that minimally acceptable control of the aircraft can be achieved even after (1) the loss of any two CAN buses or (2) the loss of any two FCCs. Furthermore, the embodiment of Figure 13A described above may offer the benefit of smaller wire weight compared to the disclosed embodiment utilizing six CAN buses. 【0100】 Additional aspects of this disclosure may be described further through the following clauses. 1. An aircraft, Torso and, One or more flight control computers configured to provide control signals, Two wings, the wings extending from both sides of the fuselage, A first set of electrically driven propellers is provided behind the wing and on both sides of the fuselage, A second set of electrically driven propellers is positioned in front of the wing and on both sides of the fuselage, The system comprises a plurality of electric buses coupled to one or more flight control computers, An aircraft in which one or more flight control computers are configured to provide control signals to only one of the first set of propellers mounted on one of the wings and only one of the second set of propellers mounted on the other wing, via the same electric bus among the plurality of electric buses. 2. The aircraft according to Clause 1, wherein the first set of propellers is tiltable between a vertical lift configuration and a forward thrust configuration. 3. The aircraft according to Clause 1 or 2, wherein the propellers of the second set are tiltable between a vertical lift configuration and a forward thrust configuration. 4. The propellers of the aforementioned first set are configured to provide vertical lift, as described in any one of the clauses 1 to 3 of the aircraft. 5. The propellers of the second set described above are configured to provide vertical lift, as described in any one of the clauses 1 to 4 of the aircraft. 6. The aforementioned multiple electric buses are The aircraft according to any one of the clauses 1 to 5, further configured such that at least one of the first set of propellers mounted furthest from the fuselage on one of the wings and at least one of the second set of propellers mounted furthest from the fuselage on the other wing are provided with control signals via the same electric bus of the plurality of electric buses. 7. The aforementioned multiple electric buses are The aircraft according to any one of the clauses 1 to 6, further configured such that at least one of the first set of propellers mounted closest to the fuselage on one of the wings and at least one of the second set of propellers mounted closest to the fuselage on the other wing are provided with control signals via the same electric bus of the plurality of electric buses. 8. The aforementioned multiple electric buses are The aircraft according to any one of the clauses 1 to 7, further configured such that control signals are provided via the same electric bus of the plurality of electric buses, the propeller of the first set of propellers mounted between the propeller of the first set of propellers mounted on one of the wings furthest from the fuselage and the propeller of the first set of propellers mounted on the same wing closest to the fuselage, and the propeller of the second set of propellers mounted between the propeller of the second set of propellers mounted on the other wing furthest from the fuselage and the propeller of the second set of propellers mounted on the same wing closest to the fuselage. 9. The aforementioned multiple electric buses are The aircraft according to any one of the clauses 1 to 8, wherein at least the propellers of the first set of propellers located on both sides of the fuselage and furthest from the fuselage are further configured to receive control signals via the same electric bus among the plurality of electric buses. 10. The aforementioned multiple electric buses are The aircraft according to any one of the clauses 1 to 9, wherein at least the propellers of the first set of propellers located on both sides of the fuselage and closest to the fuselage are further configured to receive control signals via the same electric bus of the plurality of electric buses. 11. The aforementioned multiple electric buses are The aircraft according to any one of the clauses 1 to 10, wherein at least one of the propellers in the first set of propellers, which is located on both sides and between the nearest and furthest propellers in the first set of propellers, is further configured to receive control signals via the same electric bus among the plurality of electric buses. 12. The aforementioned multiple electric buses are The aircraft according to any one of the clauses 1 to 11, wherein at least the propellers of the second set of propellers located on both sides of the fuselage and furthest from the fuselage are further configured to receive control signals via the same electric bus among the plurality of electric buses. 13. The aforementioned multiple electric buses are The aircraft according to any one of the clauses 1 to 12, wherein at least the propellers of the second set of propellers located on both sides of the fuselage and closest to the fuselage are further configured to receive control signals via the same electric bus among the plurality of electric buses. 14. The aforementioned multiple electric buses are The aircraft according to any one of the clauses 1 to 13, wherein at least one of the propellers in the second set of propellers, which is located on both sides and between the nearest and furthest propellers in the second set of propellers, is further configured to receive control signals via the same electric bus among the plurality of electric buses. 15. The aircraft according to any one of Clauses 1 to 14, further comprising a plurality of tilt propeller actuators configured to tilt the propeller between a vertical lift configuration and a forward thrust configuration, wherein the tilt propeller actuators are mounted on the wings and configured to be located on both sides of the fuselage. 16. The aforementioned multiple electric buses are The aircraft according to Clause 15, wherein at least the tilt propeller actuators located on both sides of the fuselage and furthest from the fuselage are further configured to receive control signals via the same electric bus among the plurality of electric buses. 17. The aforementioned multiple electric buses, The aircraft according to Clause 15 or 16, wherein at least the tilt propeller actuators located on both sides of the fuselage and closest to the fuselage are further configured to receive control signals via the same electric bus among the plurality of electric buses. 18. The aforementioned multiple electric buses, The aircraft according to any one of the clauses 15 to 17, wherein at least the tilt propeller actuators located on both sides and between the nearest tilt propeller actuator and the furthest tilt propeller actuator are further configured to receive control signals via the same electric bus among the plurality of electric buses. 19. An aircraft according to any one of the clauses 1 to 18, further comprising a plurality of flaperon actuators mounted on the wings and configured to be arranged on both sides of the fuselage. 20. The aforementioned multiple electric buses are The aircraft according to Clause 19, wherein at least one of the flaperon actuators mounted on one of the wings and one of the flaperon actuators mounted on the other wing are provided with a control signal via the same electric bus of the plurality of electric buses. 21. The aforementioned multiple electric buses are The aircraft according to Clause 19 or 20, wherein at least one of the flaperon actuators mounted on one of the wings and one of the flaperon actuators mounted on the other wing are provided with control signals via different electric buses among the plurality of electric buses. 22. The aforementioned multiple electric buses are The aircraft according to any one of the clauses 19 to 21, wherein at least the flaperon actuators located on both sides of the fuselage and closest to the fuselage are further configured to receive control signals via the same electric bus among the plurality of electric buses. 23. The aforementioned multiple electric buses, The aircraft according to any one of the clauses 19 to 22, wherein at least the flaperon actuators located on both sides of the fuselage and furthest from the fuselage are further configured to receive control signals via different electric buses among the plurality of electric buses. 24. An aircraft according to any one of the clauses 19 to 23, wherein the number of flaperone actuators is at least two, at least one of the flaperone actuators is mounted on one of the wings, and at least another of the flaperone actuators is mounted on the other wing. 25. An aircraft according to any one of clauses 1 to 24, further comprising a plurality of rudder actuators disposed on both sides of the fuselage. 26. The aforementioned multiple electric buses are The aircraft according to Clause 25, wherein at least the rudder actuators located on both sides of the fuselage and furthest from the fuselage are further configured to receive control signals via the same electric bus among the plurality of electric buses. 27. The aforementioned multiple electric buses are The aircraft according to Clause 25 or 26, wherein at least the rudder actuators located on both sides of the fuselage and closest to the fuselage are further configured to receive control signals via the same electric bus among the plurality of electric buses. 28. The aforementioned multiple electric buses are The aircraft according to any one of the clauses 25 to 27, wherein at least one of the rudder actuators located on both sides and between the closest rudder actuator and the furthest rudder actuator is further configured to receive control signals via the same electric bus among the plurality of electric buses. 29. An aircraft as described in any one of clauses 25 to 28, wherein the number of rudder-vator actuators is at least two, at least one of which is mounted on the first aft stabilizer, and at least another of which is mounted on the second aft stabilizer. 30. The number of propellers in the first set is six, three of the propellers in the first set are mounted on one of the wings, and the other three of the propellers in the first set are mounted on the other wing. The aircraft according to any one of Clauses 1 to 29, wherein the number of propellers in the second set is six, three of the propellers in the second set are mounted on one of the wings, and the other three of the propellers in the second set are mounted on the other wing. 31. A method for flight control, Providing control signals via at least one hardware processor included in an aircraft, the aircraft is Torso and, Two wings, the wings extending from both sides of the fuselage, A first set of electrically driven propellers is provided behind the wing and on both sides of the fuselage, A second set of electrically driven propellers is positioned in front of the wing and on both sides of the fuselage, The system includes a plurality of electric buses coupled to at least one hardware processor, Providing the control signal via the at least one hardware processor means that A method comprising providing the control signals via at least one hardware processor to one of the first set of propellers mounted on one of the wings and one of the second set of propellers mounted on the other wing, via the same electric bus among the plurality of electric buses. 32. Providing the control signal via the at least one hardware processor is: The method according to clause 31, further comprising providing the control signals via the at least one hardware processor to at least one of the first set of propellers mounted furthest from the fuselage on one of the wings and at least one of the second set of propellers mounted furthest from the fuselage on the other wing, via the same electric bus of the plurality of electric buses. 33. Providing the control signal via the at least one hardware processor is: The method according to clause 31 or 32, further comprising providing the control signals via the at least one hardware processor to at least one of the first set of propellers mounted closest to the fuselage on one of the wings, and to at least one of the second set of propellers mounted closest to the fuselage on the other wing, via the same electric bus of the plurality of electric buses. 34. Providing the control signal via the at least one hardware processor is: The method according to any one of the claims 31 to 33, further comprising providing the control signals via the at least one hardware processor to at least one of the propellers of the first set mounted between the propeller of the first set mounted furthest from the fuselage on one of the wings and the propeller of the first set mounted closest to the fuselage on the same wing, via the same electric bus of the plurality of electric buses, and at least one of the propellers of the second set mounted between the propeller of the second set mounted furthest from the fuselage on the other wing and the propeller of the second set mounted closest to the fuselage on the same wing. 35. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 31 to 34, further comprising providing the control signals via the at least one hardware processor to at least the propellers of the first set of propellers located on both sides of the fuselage and furthest from the fuselage, via the same electric bus of the plurality of electric buses. 36. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 31 to 35, further comprising providing the control signals via the at least one hardware processor to at least the propellers of the first set of propellers located on both sides of the fuselage and closest to the fuselage, via the same electric bus of the plurality of electric buses. 37. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 31 to 36, further comprising providing the control signals via at least one hardware processor to at least the propellers in the first set of propellers located on both sides and between the nearest and furthest propellers in the first set of propellers, via the same electric bus of the plurality of electric buses. 38. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 31 to 37, further comprising providing the control signals via the at least one hardware processor to at least the propellers of the second set of propellers located on both sides of the fuselage and furthest from the fuselage, via the same electric bus of the plurality of electric buses. 39. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 31 to 38, further comprising providing the control signals via the at least one hardware processor to at least the propellers in the second set of propellers located on both sides of the fuselage and closest to the fuselage, via the same electric bus of the plurality of electric buses. 40. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 31 to 39, further comprising providing the control signals via at least one hardware processor to at least the propellers in the second set of propellers, which are located on both sides and between the nearest propeller and the furthest propeller in the second set of propellers, via the same electric bus of the plurality of electric buses. 41. Providing the control signal via the at least one hardware processor is: The method according to any one of the claims 31 to 40, further comprising providing the control signals via at least one hardware processor to each propeller in the first set of propellers and each propeller in the second set of propellers via a plurality of electric buses. 42. The aforementioned aircraft, The method according to any one of claims 31 to 41, further comprising a plurality of tilt propeller actuators configured to tilt the propeller between a vertical lift configuration and a forward thrust configuration, wherein the tilt propeller actuators are mounted on the wing and configured to be located on both sides of the fuselage. 43. Providing the control signal via the at least one hardware processor is: The method according to clause 42, further comprising providing the control signals via the at least one hardware processor to at least the tilt propeller actuators located on both sides of the fuselage and furthest from the fuselage, via the same electric bus among the plurality of electric buses. 44. Providing the control signal via the at least one hardware processor is: The method according to clause 42 or 43, further comprising providing the control signals via the at least one hardware processor to at least the tilt propeller actuators located on both sides of the fuselage and closest to the fuselage, via the same electric bus among the plurality of electric buses. 45. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 42 to 44, further comprising providing the control signals via the at least one hardware processor to at least the tilt propeller actuators located on both sides of the fuselage and between the nearest tilt propeller actuator and the furthest tilt propeller actuator, via the same electric bus of the plurality of electric buses. 46. ​​The aforementioned aircraft, The method according to any one of the claims 31 to 45, further comprising a plurality of flaperon actuators mounted on the wing and configured to be arranged on both sides of the fuselage. 47. Providing the control signal via the at least one hardware processor is: The method according to clause 46, further comprising providing the control signal via the at least one hardware processor to at least one of the flaperon actuators mounted on one of the wings and one of the flaperon actuators mounted on the other wing, via the same electric bus among the plurality of electric buses. 48. Providing the control signal via the at least one hardware processor is: The method according to clause 46 or 47, further comprising providing the control signal via the at least one hardware processor to at least one of the flaperon actuators mounted on one of the wings and one of the flaperon actuators mounted on the other wing, via different electric buses among the plurality of electric buses. 49. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 46 to 48, further comprising providing the control signals via at least one hardware processor to at least the flaperon actuators located on both sides of the fuselage and furthest from the fuselage, via different electric buses among the plurality of electric buses. 50. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 46 to 49, further comprising providing the control signal via the at least one hardware processor to at least the flaperon actuators located on both sides of the fuselage and closest to the fuselage via the same electric bus among the plurality of electric buses. 51. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 46 to 50, further comprising providing the control signal via the at least one hardware processor to at least the flaperon actuators located on both sides of the fuselage and furthest from the fuselage, via the same electric bus among the plurality of electric buses. 52. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 46 to 51, further comprising providing the control signals via at least one hardware processor to at least the flaperon actuators located on both sides of the fuselage and closest to the fuselage via different electric buses among the plurality of electric buses. 53. The aforementioned aircraft, The method according to any one of the claims 31 to 52, further comprising a plurality of rudder actuators disposed on both sides of the fuselage. 54. Providing the control signal via the at least one hardware processor is: The method according to clause 53, further comprising providing the control signals via the at least one hardware processor to at least the rudder actuators located on both sides of the fuselage and furthest from the fuselage, via the same electric bus among the plurality of electric buses. 55. Providing the control signal via the at least one hardware processor is: The method according to clause 53 or 54, further comprising providing the control signals via the at least one hardware processor to at least the rudder actuators located on both sides of the fuselage and closest to the fuselage, via the same electric bus among the plurality of electric buses. 56. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 53 to 55, further comprising providing the control signals via the at least one hardware processor to at least the rudder actuators located on both sides and between the nearest rudder actuator and the furthest rudder actuator, via the same electric bus among the plurality of electric buses. 57. Aircraft, Torso and, One or more flight control computers configured to provide control signals, A first set of electrically driven propellers, and a second set of electrically driven propellers disposed on one side of the fuselage, wherein the first set comprises the first set of electrically driven propellers and the second set of electrically driven propellers, disposed in front of the second set. A third set of electrically driven propellers, and a fourth set of electrically driven propellers arranged on the other side of the fuselage, wherein the third set is arranged in front of the fourth set, and comprises the third set of electrically driven propellers and the fourth set of electrically driven propellers. The system comprises a plurality of electric buses coupled to one or more flight control computers, The one or more flight control computers are configured to provide control signals to at least one of the first set of propellers and at least one of the fourth set of propellers via one of the multiple electric buses. An aircraft in which one or more flight control computers are configured to provide control signals to at least one of the third set of propellers and at least one of the second set of propellers via another electric bus among the plurality of electric buses. 58. The aircraft according to Clause 57, wherein the first set of propellers and the third set of propellers are tiltable between a vertical lift configuration and a forward thrust configuration. 59. The aircraft according to Clause 57 or 58, wherein the first set of propellers and the third set of propellers are configured to provide vertical lift. 60. An aircraft according to any one of the clauses 57 to 59, wherein the second set of propellers and the fourth set of propellers are tiltable between a vertical lift configuration and a forward thrust configuration. 61. An aircraft according to any one of the clauses 57 to 60, wherein the second set of propellers and the fourth set of propellers are configured to provide vertical lift. 62. An aircraft according to any one of clauses 57 to 61, wherein one or more flight control computers are configured to provide control signals to one of the first set of propellers and one of the fourth set of propellers, both of which are located furthest from the fuselage, via the same electric bus of the plurality of electric buses. 63. An aircraft according to any one of clauses 57 to 62, wherein one or more flight control computers are configured to provide control signals to one of the first set of propellers and one of the fourth set of propellers, both located closest to the fuselage, via the same electric bus of the plurality of electric buses. 64. An aircraft according to any one of the clauses 57 to 63, wherein one or more flight control computers are configured to provide control signals to one of the first set of propellers, both located between the nearest and furthest propellers, and one of the fourth set of propellers, via the same electric bus of the plurality of electric buses. 65. An aircraft according to any one of clauses 57 to 64, wherein one or more flight control computers are configured to provide control signals to one of the third set of propellers and one of the second set of propellers, both of which are located furthest from the fuselage, via the same electric bus of the plurality of electric buses. 66. An aircraft according to any one of clauses 57 to 65, wherein one or more flight control computers are configured to provide control signals to one of the third set of propellers and one of the second set of propellers, both of which are located closest to the fuselage, via the same electric bus of the plurality of electric buses. 67. An aircraft according to any one of the clauses 57 to 66, wherein one or more flight control computers are configured to provide control signals to one of the third set of propellers, both located between the nearest and furthest propellers, and one of the second set of propellers, via the same electric bus of the plurality of electric buses. 68. The aircraft according to any one of the clauses 57 to 67, wherein one or more flight control computers are configured to provide control signals to at least one of the first set of propellers and at least one of the third set of propellers via the same electric bus among the plurality of electric buses. 69. An aircraft according to any one of the clauses 57 to 68, wherein one or more flight control computers are configured to provide control signals to one of the first set of propellers and one of the third set of propellers, both of which are located furthest from the fuselage, via the same electric bus of the plurality of electric buses. 70. An aircraft according to any one of the clauses 57 to 69, wherein one or more flight control computers are configured to provide control signals to one of the first set of propellers and one of the third set of propellers, both of which are located closest to the fuselage, via the same electric bus of the plurality of electric buses. 71. The aircraft according to any one of the clauses 57 to 70, wherein one or more flight control computers are configured to provide control signals to one of the first set of propellers and one of the third set of propellers, both located between the nearest and furthest propellers, via the same electric bus of the plurality of electric buses. 72. An aircraft according to any one of the clauses 57 to 71, wherein one or more flight control computers are configured to provide control signals to at least one of the second set of propellers and at least one of the fourth set of propellers via the same electric bus among the plurality of electric buses. 73. An aircraft according to any one of Clauses 57 to 72, wherein one or more flight control computers are configured to provide control signals to one of the second set of propellers and one of the fourth set of propellers, both of which are located furthest from the fuselage, via the same electric bus of the plurality of electric buses. 74. An aircraft according to any one of clauses 57 to 73, wherein one or more flight control computers are configured to provide control signals to one of the second set of propellers and one of the fourth set of propellers, both of which are located closest to the fuselage, via the same electric bus of the plurality of electric buses. 75. An aircraft according to any one of clauses 57 to 74, wherein one or more flight control computers are configured to provide control signals to one of the second set of propellers, both located between the nearest and furthest propellers, and one of the fourth set of propellers, via the same electric bus of the plurality of electric buses. 76. An aircraft according to any one of the clauses 57 to 75, further comprising a plurality of tilt propeller actuators, each configured to tilt one of the electric propellers between a vertical lift configuration and a forward thrust configuration, wherein the tilt propeller actuators are located on both sides of the fuselage. 77. The aircraft according to Clause 76, wherein one or more flight control computers are configured to provide control signals to the tilt propeller actuators located on both sides of the fuselage and furthest from the fuselage, via the same electric bus among the plurality of electric buses. 78. The aircraft according to Clause 76 or 77, wherein one or more flight control computers are configured to provide control signals to the tilt propeller actuators located on both sides of the fuselage and closest to the fuselage, via the same electric bus of the plurality of electric buses. 79. The aircraft according to any one of the clauses 76 to 78, wherein one or more flight control computers are configured to provide control signals to the tilt propeller actuators located on both sides and between the closest tilt propeller actuator and the furthest tilt propeller actuator via the same electric bus of the plurality of electric buses. 80. An aircraft according to any one of the clauses 57 to 79, further comprising a plurality of flaperon actuators disposed on both sides of the fuselage. 81. The aircraft according to Clause 80, wherein one or more flight control computers are configured to provide control signals to the flaperon actuators located on both sides of the fuselage and furthest from the fuselage, via the same electric bus among the plurality of electric buses. 82. The aircraft according to Clause 80 or 81, wherein one or more flight control computers are configured to provide control signals to the flaperon actuators located on both sides of the fuselage and closest to the fuselage via the same electric bus of the plurality of electric buses. 83. The aircraft according to any one of the clauses 80 to 82, wherein one or more flight control computers are configured to provide control signals to the flaperon actuators located on both sides of the fuselage and furthest from the fuselage via different electric buses among the plurality of electric buses. 84. The aircraft according to any one of the clauses 80 to 83, wherein one or more flight control computers are configured to provide control signals to the flaperon actuators located on both sides of the fuselage and closest to the fuselage via different electric buses among the plurality of electric buses. 85. The aircraft according to any one of the clauses 80 to 84, wherein the number of flaperon actuators is at least two, at least one of the flaperon actuators is located on one side of the fuselage, and at least another of the flaperon actuators is located on the other side of the fuselage. 86. An aircraft according to any one of clauses 57 to 85, further comprising a plurality of rudder actuators disposed on both sides of the fuselage. 87. The aircraft according to Clause 86, wherein one or more flight control computers are configured to provide control signals to the rudder actuators located on both sides of the fuselage and furthest from the fuselage via the same electric bus among the plurality of electric buses. 88. The aircraft according to Clause 86 or 87, wherein one or more flight control computers are configured to provide control signals to the rudder actuators located on both sides of the fuselage and closest to the fuselage via the same electric bus of the plurality of electric buses. 89. The aircraft according to any one of the clauses 86 to 88, wherein one or more flight control computers are configured to provide control signals to the rudder actuators located on both sides and between the closest rudder actuator and the furthest rudder actuator via the same electric bus of the plurality of electric buses. 90. The aircraft according to any one of the clauses 86 to 89, wherein the number of rudder actuators is at least two, at least one of the rudder actuators is located on one side of the fuselage, and at least another of the rudder actuators is located on the other side of the fuselage. 91. The number of propellers in each set of propellers is three, as specified in any one of clauses 57-90. 92. The number of propellers in each set of propellers is two, as specified in any one of clauses 57-90. 93. An aircraft according to any one of clauses 57 to 92, wherein one or more flight control computers are configured to provide control signals to several devices located on one side of the fuselage and an equal number of devices located on the other side of the fuselage via the same electric bus among the plurality of electric buses. 94. A method for flight control, Providing control signals via at least one hardware processor included in an aircraft, the aircraft is Torso and, One or more flight control computers configured to provide control signals, A first set of electrically driven propellers, and a second set of electrically driven propellers disposed on one side of the fuselage, wherein the first set comprises the first set of electrically driven propellers and the second set of electrically driven propellers, disposed in front of the second set. A third set of electrically driven propellers, and a fourth set of electrically driven propellers arranged on the other side of the fuselage, wherein the third set is arranged in front of the fourth set, and comprises the third set of electrically driven propellers and the fourth set of electrically driven propellers. The system comprises a plurality of electric buses coupled to one or more flight control computers, The one or more flight control computers are configured to provide control signals to at least one of the first set of propellers and at least one of the fourth set of propellers via one of the multiple electric buses. A method wherein one or more flight control computers are configured to provide control signals to at least one of the third set of propellers and at least one of the second set of propellers via another electric bus among the plurality of electric buses. 95. The method according to Clause 94, wherein the first set of propellers and the third set of propellers are tiltable between a vertical lift configuration and a forward thrust configuration. 96. The method according to clause 94 or 95, wherein the first set of propellers and the third set of propellers are configured to provide vertical lift. 97. The method according to any one of the clauses 94 to 96, wherein the first set of propellers and the third set of propellers are tiltable between a vertical lift configuration and a forward thrust configuration. 98. The method according to any one of the clauses 94 to 97, wherein the second set of propellers and the fourth set of propellers are tiltable between a vertical lift configuration and a forward thrust configuration. 99. The method according to any one of the clauses 94 to 98, wherein the second set of propellers and the fourth set of propellers are configured to provide vertical lift. 100. The method according to any one of the clauses 94 to 99, wherein the second set of propellers and the fourth set of propellers are tiltable between a vertical lift configuration and a forward thrust configuration. 101. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 94 to 100, further comprising providing the control signals via at least one hardware processor to one of the first set of propellers and one of the fourth set of propellers, both of which are located furthest from the fuselage, via the same electric bus of the plurality of electric buses. 102. Providing the control signal via the at least one hardware processor is: The method according to any one of the claims 94 to 101, further comprising providing the control signals via at least one hardware processor to one of the first set of propellers and one of the fourth set of propellers, both located closest to the fuselage, via the same electric bus of the plurality of electric buses. 103. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 94 to 102, further comprising providing the control signals via at least one hardware processor to one of the first set of propellers and one of the fourth set of propellers, both located between the nearest and furthest propellers, via the same electric bus of the plurality of electric buses. 104. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 94 to 103, further comprising providing the control signals via at least one hardware processor to one of the third set of propellers and one of the second set of propellers, both of which are located furthest from the fuselage, via the same electric bus of the plurality of electric buses. 105. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 94 to 104, further comprising providing the control signals via at least one hardware processor to one of the third set of propellers and one of the second set of propellers, both located closest to the fuselage, via the same electric bus of the plurality of electric buses. 106. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 94 to 105, further comprising providing the control signals via at least one hardware processor to one of the third set of propellers and one of the second set of propellers, both located between the nearest and furthest propellers, via the same electric bus of the plurality of electric buses. 107. Providing the control signal via the at least one hardware processor is: The method according to any one of the claims 94 to 106, further comprising providing the control signals via the at least one hardware processor to at least one of the first set of propellers and at least one of the third set of propellers via the same electric bus among the plurality of electric buses. 108. Providing the control signal via the at least one hardware processor is: The method according to any one of the claims 94 to 107, further comprising providing the control signals via at least one hardware processor to one of the first set of propellers and one of the third set of propellers, both located furthest from the fuselage, via the same electric bus of the plurality of electric buses. 109. Providing the control signal via the at least one hardware processor is: The method according to any one of the claims 94 to 108, further comprising providing the control signals via at least one hardware processor to one of the first set of propellers and one of the third set of propellers, both located closest to the fuselage, via the same electric bus of the plurality of electric buses. 110. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 94 to 109, further comprising providing the control signals via at least one hardware processor to one of the first set of propellers and one of the third set of propellers, both located between the nearest and furthest propellers, via the same electric bus of the plurality of electric buses. 111. Providing the control signal via the at least one hardware processor is: The method according to any one of the claims 94 to 110, further comprising providing the control signals via the at least one hardware processor to at least one of the second set of propellers and at least one of the fourth set of propellers via the same electric bus among the plurality of electric buses. 112. Providing the control signal via the at least one hardware processor is: The method according to any one of the claims 94 to 111, further comprising providing the control signals via at least one hardware processor to one of the second set of propellers and one of the fourth set of propellers, both of which are located furthest from the fuselage, via the same electric bus of the plurality of electric buses. 113. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 94 to 112, further comprising providing the control signals via at least one hardware processor to one of the second set of propellers and one of the fourth set of propellers, both located closest to the fuselage, via the same electric bus of the plurality of electric buses. 114. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 94 to 113, further comprising providing the control signals via at least one hardware processor to one of the second set of propellers and one of the fourth set of propellers, both located between the nearest and furthest propellers, via the same electric bus of the plurality of electric buses. 115. The aforementioned aircraft, The method according to any one of the claims 94 to 114, further comprising a plurality of tilt propeller actuators, each configured to tilt one of the electric propellers between a vertical lift configuration and a forward thrust configuration, wherein the tilt propeller actuators are located on both sides of the fuselage. 116. Providing the control signal via the at least one hardware processor is: The method according to clause 115, further comprising providing the control signals via at least one hardware processor to the tilt propeller actuators located on both sides of the fuselage and furthest from the fuselage, via the same electric bus among the plurality of electric buses. 117. Providing the control signal via the at least one hardware processor is: The method according to clause 115 or 116, further comprising providing the control signals via at least one hardware processor to the tilt propeller actuators located on both sides of the fuselage and closest to the fuselage via the same electric bus among the plurality of electric buses. 118. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 115 to 117, further comprising providing the control signals via at least one hardware processor to the tilt propeller actuators located on both sides and between the closest tilt propeller actuator and the furthest tilt propeller actuator, via the same electric bus among the plurality of electric buses. 119. The aforementioned aircraft, The method according to any one of the claims 94 to 118, further comprising a plurality of flaperon actuators configured to be disposed on both sides of the fuselage. 120. Providing the control signal via the at least one hardware processor is: The method according to clause 119, further comprising providing the control signals via at least one hardware processor to the flaperon actuators located on both sides of the fuselage and furthest from the fuselage, via the same electric bus among the plurality of electric buses. 121. Providing the control signal via the at least one hardware processor is: The method according to clause 119 or 120, further comprising providing the control signals via at least one hardware processor to the flaperon actuators located on both sides of the fuselage and closest to the fuselage via the same electric bus among the plurality of electric buses. 122. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 119 to 121, further comprising providing the control signals via at least one hardware processor to the flaperon actuators located on both sides of the fuselage and furthest from the fuselage, via different electric buses among the plurality of electric buses. 123. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 119 to 122, further comprising providing the control signals via at least one hardware processor to the flaperon actuators located on both sides of the fuselage and closest to the fuselage via different electric buses among the plurality of electric buses. 124. The method according to any one of claims 119 to 123, wherein the number of flaperone actuators is at least two, at least one of the flaperone actuators is disposed on one side of the body, and at least another of the flaperone actuators is disposed on the other side of the body. 125. The aforementioned aircraft, The method according to any one of the claims 94 to 124, further comprising a plurality of rudder actuators configured to be disposed on both sides of the fuselage. 126. Providing the control signal via the at least one hardware processor is: The method according to clause 125, further comprising providing the control signals via at least one hardware processor to the rudder actuators located on both sides of the fuselage and furthest from the fuselage, via the same electric bus among the plurality of electric buses. 127. Providing the control signal via the at least one hardware processor is: The method according to clause 125 or 126, further comprising providing the control signals via at least one hardware processor to the rudder actuators located on both sides of the fuselage and closest to the fuselage via the same electric bus among the plurality of electric buses. 128. Providing the control signal via the at least one hardware processor is: The method according to any one of claims 125 to 127, further comprising providing the control signals via at least one hardware processor to the rudder actuators located on both sides and between the closest rudder actuator and the furthest rudder actuator, via the same electric bus among the plurality of electric buses. 129. The method according to any one of the claims 125 to 128, wherein the number of rudder-vator actuators is at least two, at least one of the rudder-vator actuators is located on one side of the fuselage, and at least another of the rudder-vator actuators is located on the other side of the fuselage. 130. The number of propellers in each set of propellers is three, as described in any one of clauses 94 to 129. 131. The number of propellers in each set of propellers is two, as described in any one of the clauses 94 to 129. 132. Providing the control signal via the at least one hardware processor is: The method according to any one of the claims 94 to 131, further comprising providing the control signals via at least one hardware processor to several devices located on one side of the fuselage and an equal number of devices located on the other side of the fuselage via the same electric bus among the plurality of electric buses. 【0101】 The above description is provided for illustrative purposes only. It is not exhaustive and does not limit the invention to the exact form or embodiment disclosed herein. Modifications and adaptations of the invention will be apparent to those skilled in the art from the examination herein and the practice of the disclosed embodiments of the invention disclosed herein. 【0102】 The features and advantages of this disclosure are evident from the detailed specification, and therefore the attached claims are intended to cover all systems and methods included in the true spirit and scope of this disclosure. As used herein, the indefinite articles "a" and "an" mean "one or more." Similarly, the use of plural terms does not necessarily indicate plural unless it is clear in the given context. Words such as "and" or "or" mean "and / or" unless otherwise indicated. Furthermore, since numerous modifications and variations can easily arise from examining this disclosure, it is not desirable to limit this disclosure to the exact configurations and operations illustrated and described, and therefore all suitable modifications and equivalents included in the scope of this disclosure may be applicable. 【0103】 Other embodiments will be apparent to those skilled in the art from the examination of this specification and the practice of the implementations disclosed herein. The architectures and circuit layouts shown in the figures are intended for illustrative purposes only and are not intended to be limited to the specific configurations and circuit layouts shown in the figures. Furthermore, this specification and the examples are intended to be considered merely illustrative, and the true scope and spirit of the invention are set forth by the following claims. The above description is presented for illustrative purposes only. It is not exhaustive and does not limit the invention to the exact forms or embodiments disclosed herein. Modifications and adaptations of the invention will be apparent to those skilled in the art from the examination of this specification and the practice of the disclosed embodiments of the invention disclosed herein.

Claims

[Claim 1] It is an aircraft, Torso and, One or more flight control computers configured to provide control signals, A first set of electric propellers and a second set of electric propellers are arranged on one side of the fuselage, wherein the first set is arranged in front of the second set, and comprises the first set of electric propellers and the second set of electric propellers. A third set of electric propellers and a fourth set of electric propellers are arranged on another side of the fuselage, wherein the third set is arranged in front of the fourth set, and comprises the third set of electric propellers and the fourth set of electric propellers. The system comprises a plurality of electric buses coupled to one or more flight control computers, The one or more flight control computers are configured to provide control signals to at least one of the first set of propellers and at least one of the fourth set of propellers via one of the multiple electric buses. An aircraft in which one or more flight control computers are configured to provide control signals to at least one of the third set of propellers and at least one of the second set of propellers via another electric bus among the plurality of electric buses. [Claim 2] The aircraft according to claim 1, wherein the first set of propellers and the third set of propellers are tiltable between a vertical lift configuration and a forward thrust configuration. [Claim 3] The aircraft according to claim 1 or 2, wherein the first set of propellers and the third set of propellers are configured to provide vertical lift. [Claim 4] The aircraft according to any one of claims 1 to 3, wherein the second set of propellers and the fourth set of propellers are tiltable between a vertical lift configuration and a forward thrust configuration. [Claim 5] The aircraft according to any one of claims 1 to 4, wherein the second set of propellers and the fourth set of propellers are configured to provide vertical lift. [Claim 6] The aircraft according to any one of claims 1 to 5, wherein one or more flight control computers are configured to provide control signals to one of the first set of propellers and one of the fourth set of propellers, both of which are located furthest from the fuselage, via the same electric bus among the plurality of electric buses. [Claim 7] The aircraft according to any one of claims 1 to 6, wherein one or more flight control computers are configured to provide control signals to one of the first set of propellers and one of the fourth set of propellers, both of which are located closest to the fuselage, via the same electric bus among the plurality of electric buses. [Claim 8] The aircraft according to any one of claims 1 to 7, wherein one or more flight control computers are configured to provide control signals to one of the first set of propellers and one of the fourth set of propellers, both of which are located between the nearest and furthest propellers, via the same electric bus among the plurality of electric buses. [Claim 9] The aircraft according to any one of claims 1 to 8, wherein one or more flight control computers are configured to provide control signals to one of the third set of propellers and one of the second set of propellers, both of which are located furthest from the fuselage, via the same electric bus among the plurality of electric buses. [Claim 10] The aircraft according to any one of claims 1 to 9, wherein one or more flight control computers are configured to provide control signals to one of the third set of propellers and one of the second set of propellers, both of which are located closest to the fuselage, via the same electric bus among the plurality of electric buses. [Claim 11] The aircraft according to any one of claims 1 to 10, wherein one or more flight control computers are configured to provide control signals to one of the third set of propellers, both located between the nearest propeller and the furthest propeller, and one of the second set of propellers, via the same electric bus among the plurality of electric buses. [Claim 12] The aircraft according to any one of claims 1 to 11, wherein one or more flight control computers are configured to provide control signals to at least one of the first set of propellers and at least one of the third set of propellers via the same electric bus among the plurality of electric buses. [Claim 13] The aircraft according to any one of claims 1 to 12, wherein one or more flight control computers are configured to provide control signals to one of the first set of propellers and one of the third set of propellers, both of which are located furthest from the fuselage, via the same electric bus among the plurality of electric buses. [Claim 14] The aircraft according to any one of claims 1 to 13, wherein one or more flight control computers are configured to provide control signals to one of the first set of propellers and one of the third set of propellers, both of which are located closest to the fuselage, via the same electric bus among the plurality of electric buses. [Claim 15] The aircraft according to any one of claims 1 to 14, wherein one or more flight control computers are configured to provide control signals to one of the first set of propellers, both located between the nearest and furthest propellers, and one of the third set of propellers, via the same electric bus among the plurality of electric buses. [Claim 16] The aircraft according to any one of claims 1 to 15, wherein one or more flight control computers are configured to provide control signals to at least one of the second set of propellers and at least one of the fourth set of propellers via the same electric bus among the plurality of electric buses. [Claim 17] The aircraft according to any one of claims 1 to 16, wherein one or more flight control computers are configured to provide control signals to one of the second set of propellers and one of the fourth set of propellers, both of which are located furthest from the fuselage, via the same electric bus among the plurality of electric buses. [Claim 18] The aircraft according to any one of claims 1 to 17, wherein one or more flight control computers are configured to provide control signals to one of the second set of propellers and one of the fourth set of propellers, both of which are located closest to the fuselage, via the same electric bus among the plurality of electric buses. [Claim 19] The aircraft according to any one of claims 1 to 18, wherein one or more flight control computers are configured to provide control signals to one of the second set of propellers, both located between the nearest propeller and the furthest propeller, and one of the fourth set of propellers, via the same electric bus among the plurality of electric buses. [Claim 20] The aircraft according to any one of claims 1 to 19, further comprising a plurality of tilt propeller actuators, each configured to tilt one of the electric propellers between a vertical lift configuration and a forward thrust configuration, wherein the tilt propeller actuators are disposed on both sides of the fuselage. [Claim 21] The aircraft according to claim 20, wherein one or more flight control computers are configured to provide control signals to the tilt propeller actuators located on both sides of the fuselage and furthest from the fuselage, via the same electric bus among the plurality of electric buses. [Claim 22] The aircraft according to claim 20 or 21, wherein one or more flight control computers are configured to provide control signals to the tilt propeller actuators located on both sides of the fuselage and closest to the fuselage, via the same electric bus among the plurality of electric buses. [Claim 23] The aircraft according to any one of claims 20 to 22, wherein one or more flight control computers are configured to provide control signals to the tilt propeller actuators located on both sides and between the closest tilt propeller actuator and the furthest tilt propeller actuator via the same electric bus among the plurality of electric buses. [Claim 24] The aircraft according to any one of claims 1 to 23, further comprising a plurality of flaperon actuators disposed on both sides of the fuselage. [Claim 25] The aircraft according to claim 24, wherein one or more flight control computers are configured to provide control signals to the flaperon actuators located on both sides of the fuselage and furthest from the fuselage via the same electric bus among the plurality of electric buses. [Claim 26] The aircraft according to claim 24 or 25, wherein one or more flight control computers are configured to provide control signals to the flaperon actuators located on both sides of the fuselage and closest to the fuselage via the same electric bus among the plurality of electric buses. [Claim 27] The aircraft according to any one of claims 24 to 26, wherein one or more flight control computers are configured to provide control signals to the flaperon actuators located on both sides of the fuselage and furthest from the fuselage via different electric buses among the plurality of electric buses. [Claim 28] The aircraft according to any one of claims 24 to 27, wherein one or more flight control computers are configured to provide control signals to the flaperon actuators located on both sides of the fuselage and closest to the fuselage via different electric buses among the plurality of electric buses. [Claim 29] The aircraft according to any one of claims 24 to 28, wherein the number of flaperone actuators is at least two, at least one of the flaperone actuators is disposed on one side of the fuselage, and at least another of the flaperone actuators is disposed on the other side of the fuselage. [Claim 30] The aircraft according to any one of claims 1 to 29, further comprising a plurality of rudder actuators disposed on both sides of the fuselage. [Claim 31] The aircraft according to claim 30, wherein one or more flight control computers are configured to provide control signals to the rudder actuators located on both sides of the fuselage and furthest from the fuselage via the same electric bus among the plurality of electric buses. [Claim 32] The aircraft according to claim 30 or 31, wherein one or more flight control computers are configured to provide control signals to the rudder actuators located on both sides of the fuselage and closest to the fuselage via the same electric bus among the plurality of electric buses. [Claim 33] The aircraft according to any one of claims 30 to 32, wherein one or more flight control computers are configured to provide control signals to the rudder actuators located on both sides and between the closest rudder actuator and the furthest rudder actuator via the same electric bus among the plurality of electric buses. [Claim 34] The aircraft according to any one of claims 30 to 33, wherein the number of rudder actuators is at least two, at least one of the rudder actuators is located on one side of the fuselage, and at least another of the rudder actuators is located on the other side of the fuselage. [Claim 35] The aircraft according to any one of claims 1 to 34, wherein the number of propellers in each set of propellers is three. [Claim 36] The aircraft according to any one of claims 1 to 35, wherein the number of propellers in each set of propellers is two. [Claim 37] The aircraft according to any one of claims 1 to 36, wherein one or more flight control computers are configured to provide control signals to several devices located on one side of the fuselage and an equal number of devices located on the other side of the fuselage via the same electric bus among the plurality of electric buses. [Claim 38] A method for flight control, Providing control signals via at least one hardware processor included in an aircraft, the aircraft is Torso and, One or more flight control computers configured to provide control signals, A first set of electrically driven propellers, and a second set of electrically driven propellers disposed on one side of the fuselage, wherein the first set comprises the first set of electrically driven propellers and the second set of electrically driven propellers, disposed in front of the second set. A third set of electrically driven propellers, and a fourth set of electrically driven propellers disposed on the other side of the fuselage, wherein the third set is disposed in front of the fourth set, and comprises the third set of electrically driven propellers and the fourth set of electrically driven propellers. The system comprises a plurality of electric buses coupled to one or more flight control computers, The one or more flight control computers are configured to provide control signals to at least one of the first set of propellers and at least one of the fourth set of propellers via one of the multiple electric buses. A method wherein one or more flight control computers are configured to provide control signals to at least one of the third set of propellers and at least one of the second set of propellers via another electric bus among the plurality of electric buses. [Claim 39] The method according to claim 38, wherein the first set of propellers and the third set of propellers are tiltable between a vertical lift configuration and a forward thrust configuration. [Claim 40] The method according to claim 38 or 39, wherein the first set of propellers and the third set of propellers are configured to provide vertical lift. [Claim 41] The method according to any one of claims 38 to 40, wherein the first set of propellers and the third set of propellers are tiltable between a vertical lift configuration and a forward thrust configuration. [Claim 42] The method according to any one of claims 38 to 41, wherein the second set of propellers and the fourth set of propellers are tiltable between a vertical lift configuration and a forward thrust configuration. [Claim 43] The method according to any one of claims 38 to 42, wherein the second set of propellers and the fourth set of propellers are configured to provide vertical lift. [Claim 44] The method according to any one of claims 38 to 43, wherein the second set of propellers and the fourth set of propellers are tiltable between a vertical lift configuration and a forward thrust configuration. [Claim 45] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 44, further comprising providing the control signals via at least one hardware processor to one of the first set of propellers and one of the fourth set of propellers, both of which are located furthest from the fuselage, via the same electric bus of the plurality of electric buses. [Claim 46] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 45, further comprising providing the control signals via at least one hardware processor to one of the first set of propellers and one of the fourth set of propellers, both located closest to the fuselage, via the same electric bus of the plurality of electric buses. [Claim 47] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 46, further comprising providing the control signals via at least one hardware processor to one of the first set of propellers and one of the fourth set of propellers, both located between the nearest and furthest propellers, via the same electric bus of the plurality of electric buses. [Claim 48] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 47, further comprising providing the control signals via at least one hardware processor to one of the third set of propellers and one of the second set of propellers, both of which are located furthest from the fuselage, via the same electric bus of the plurality of electric buses. [Claim 49] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 48, further comprising providing the control signals via at least one hardware processor to one of the third set of propellers and one of the second set of propellers, both located closest to the fuselage, via the same electric bus of the plurality of electric buses. [Claim 50] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 49, further comprising providing the control signals via at least one hardware processor to one of the third set of propellers and one of the second set of propellers, both located between the nearest and furthest propellers, via the same electric bus of the plurality of electric buses. [Claim 51] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 50, further comprising providing the control signals via the at least one hardware processor to at least one of the first set of propellers and at least one of the third set of propellers via the same electric bus among the plurality of electric buses. [Claim 52] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 51, further comprising providing the control signals via at least one hardware processor to one of the first set of propellers and one of the third set of propellers, both of which are located furthest from the fuselage, via the same electric bus of the plurality of electric buses. [Claim 53] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 52, further comprising providing the control signals via at least one hardware processor to one of the first set of propellers and one of the third set of propellers, both located closest to the fuselage, via the same electric bus among the plurality of electric buses. [Claim 54] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 53, further comprising providing the control signals via at least one hardware processor to one of the first set of propellers and one of the third set of propellers, both located between the nearest and furthest propellers, via the same electric bus of the plurality of electric buses. [Claim 55] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 54, further comprising providing the control signals via the at least one hardware processor to at least one of the second set of propellers and at least one of the fourth set of propellers via the same electric bus among the plurality of electric buses. [Claim 56] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 55, further comprising providing the control signals via at least one hardware processor to one of the second set of propellers and one of the fourth set of propellers, both of which are located furthest from the fuselage, via the same electric bus of the plurality of electric buses. [Claim 57] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 56, further comprising providing the control signals via at least one hardware processor to one of the second set of propellers and one of the fourth set of propellers, both located closest to the fuselage, via the same electric bus of the plurality of electric buses. [Claim 58] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 57, further comprising providing the control signals via at least one hardware processor to one of the second set of propellers and one of the fourth set of propellers, both located between the nearest and furthest propellers, via the same electric bus of the plurality of electric buses. [Claim 59] The aforementioned aircraft, The method according to any one of claims 38 to 58, further comprising a plurality of tilt propeller actuators, each configured to tilt one of the electric propellers between a vertical lift configuration and a forward thrust configuration, wherein the tilt propeller actuators are disposed on both sides of the fuselage. [Claim 60] Providing the control signal via the at least one hardware processor means that The method according to claim 59, further comprising providing the control signals via at least one hardware processor to the tilt propeller actuators located on both sides of the fuselage and furthest from the fuselage, via the same electric bus among the plurality of electric buses. [Claim 61] Providing the control signal via the at least one hardware processor means that The method according to claim 59 or 60, further comprising providing the control signals via at least one hardware processor to the tilt propeller actuators located on both sides of the fuselage and closest to the fuselage via the same electric bus among the plurality of electric buses. [Claim 62] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 59 to 61, further comprising providing the control signal via at least one hardware processor to the tilt propeller actuators located on both sides and between the closest tilt propeller actuator and the furthest tilt propeller actuator, via the same electric bus among the plurality of electric buses. [Claim 63] The aforementioned aircraft, The method according to any one of claims 38 to 62, further comprising a plurality of flaperon actuators configured to be arranged on both sides of the fuselage. [Claim 64] Providing the control signal via the at least one hardware processor means that The method according to claim 63, further comprising providing the control signal via at least one hardware processor to the flaperon actuators located on both sides of the fuselage and furthest from the fuselage, via the same electric bus among the plurality of electric buses. [Claim 65] Providing the control signal via the at least one hardware processor means that The method according to claim 63 or 64, further comprising providing the control signal via at least one hardware processor to the flaperon actuators located on both sides of the fuselage and closest to the fuselage via the same electric bus among the plurality of electric buses. [Claim 66] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 63 to 65, further comprising providing the control signal via at least one hardware processor to the flaperon actuators located on both sides of the fuselage and furthest from the fuselage via different electric buses among the plurality of electric buses. [Claim 67] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 63 to 66, further comprising providing the control signal via at least one hardware processor to the flaperon actuators located on both sides of the fuselage and closest to the fuselage via different electric buses among the plurality of electric buses. [Claim 68] The method according to any one of claims 63 to 67, wherein the number of flaperone actuators is at least two, at least one of the flaperone actuators is disposed on one side of the body, and at least another of the flaperone actuators is disposed on the other side of the body. [Claim 69] The aforementioned aircraft, The method according to any one of claims 38 to 68, further comprising a plurality of rudder actuators configured to be arranged on both sides of the fuselage. [Claim 70] Providing the control signal via the at least one hardware processor means that The method according to claim 69, further comprising providing the control signals via at least one hardware processor to the rudder actuators located on both sides of the fuselage and furthest from the fuselage, via the same electric bus among the plurality of electric buses. [Claim 71] Providing the control signal via the at least one hardware processor means that The method according to claim 69 or 70, further comprising providing the control signals via at least one hardware processor to the rudder actuators located on both sides of the fuselage and closest to the fuselage via the same electric bus among the plurality of electric buses. [Claim 72] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 69 to 71, further comprising providing the control signals via at least one hardware processor to the rudder actuators located on both sides and between the closest rudder actuator and the furthest rudder actuator, via the same electric bus among the plurality of electric buses. [Claim 73] The method according to any one of claims 69 to 72, wherein the number of rudder-vator actuators is at least two, at least one of the rudder-vator actuators is located on one side of the fuselage, and at least another of the rudder-vator actuators is located on the other side of the fuselage. [Claim 74] The method according to any one of claims 38 to 73, wherein the number of propellers in each set of propellers is three. [Claim 75] The method according to any one of claims 38 to 73, wherein the number of propellers in each set of propellers is two. [Claim 76] Providing the control signal via the at least one hardware processor means that The method according to any one of claims 38 to 75, further comprising providing the control signals via at least one hardware processor to several devices located on one side of the fuselage and an equal number of devices located on the other side of the fuselage via the same electric bus among the plurality of electric buses.

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