Systems and methods for controlling aircraft subsystems in different modes and states

Abstract

A method of controlling power distribution in an aircraft, the method comprising: receiving, at a control circuit in an aircraft, a selection of one of at least three aircraft operating modes from a user input device; and controlling, via the control circuit, power distribution within the aircraft based on the selected operating mode, wherein controlling power distribution based on the selected operating mode comprises separately controlling, via the control circuit, high voltage power to at least one electrically powered propulsion unit and high voltage power to at least one non-propulsion load based on the selected operating mode.

Description

Cross-references to related applications

[0001] This disclosure claims U.S. Provisional Application No. 63 / 608,107, entitled “AIRCRAFT MODES AND STATES”, filed December 8, 2023 (previous agent file number 16163.6029-00000 and current agent file number 16499.6010-00000), entitled “AIRCRAFT MODES AND STATES”; and U.S. Provisional Application No. 63 / 616,316, entitled “BATTERY MANAGEMENT SYSTEM”, filed December 29, 2023 (agent file number 16497.6005-00000), entitled “BATTERY MANAGEMENT SYSTEM”; and U.S. Provisional Application No. 63 / 616,316, entitled “BATTERY MANAGEMENT SYSTEM”, filed November 14, 2023, entitled “High Voltage Battery”. The PCT application for “Architecture”, PCT / US23 / 79690 (Agent's File No. 16163.0018-00304), which in turn claims priority and interest in U.S. Provisional Application No. 63 / 383,660 (Agent's File No. 16163.6005-00000), filed November 14, 2022, entitled “Systems and Methods for Improved Battery Assemblies for eVTOL Aircraft.” The entire contents of the foregoing applications are incorporated herein by reference for all purposes. Technical Field

[0002] This disclosure generally relates to the field of powered aircraft. More particularly, but not limited to, this disclosure relates to innovations in aircraft using electric propulsion systems. Certain aspects of this disclosure generally relate to switching between different aircraft modes based on user selection. Other aspects of this disclosure generally relate to switching between different aircraft states within a set mode based on the detection of certain events. Background Technology

[0003] Here, the inventors have recognized several issues that may be associated with the subsystems controlling aircraft, including those using electric or hybrid electric propulsion systems (hereinafter referred to as electric propulsion units or "EPUs"). For example, aircraft power supplies (e.g., battery packs) and circuitry systems need to be configured to provide redundancy, fault tolerance, energy efficiency, and weight efficiency. Additionally, users may wish to switch aircraft operating modes or states without turning various subsystems on and off. When switching to different modes or states, it is desirable to deactivate unnecessary components to conserve energy and maintain component lifespan and integrity. Summary of the Invention

[0004] This disclosure generally relates to the configuration and control of aircraft subsystems. More particularly, but not limited to, this disclosure relates to innovations in aircraft using electric or hybrid electric propulsion systems. Certain aspects of this disclosure relate to circuitry and component configurations that provide redundancy, fault tolerance, energy efficiency, and weight efficiency. Other aspects of this disclosure relate to safely controlling aircraft subsystems by confirming that certain conditions are met before switching aircraft modes. Still other aspects of this disclosure relate to control sequences and circuitry configurations that allow subsystems to be enabled and disabled while maintaining power supply to critical subsystems.

[0005] One aspect of this disclosure relates to a method for controlling power distribution in an aircraft, the method comprising: receiving, at a control circuit in the aircraft, a selection of one of at least three aircraft operating modes from a user input device; and, based on the selected operating mode, controlling power distribution within the aircraft via the control circuit, wherein controlling power distribution based on the selected operating mode includes: based on the selected operating mode, separately controlling high-voltage power supplied to at least one electric propulsion unit and high-voltage power supplied to at least one non-propulsion load via the control circuit.

[0006] Another aspect of this disclosure relates to a system for an aircraft, the system comprising: a battery management unit; one or more battery cells configured to supply high-voltage power; at least one first switching device configured to enable and disable power supply from a charging port to the one or more battery cells; at least one second switching device configured to enable and disable power supply from the one or more battery cells to a non-propulsion load; and at least one third switching device configured to enable and disable power supply from the one or more battery cells to the electric propulsion unit of the aircraft. The battery management unit controls the first, second, and third switching devices based on a selection of at least three aircraft operating modes received from a user input device.

[0007] Another aspect of this disclosure relates to a system for an aircraft, the system comprising: a user input device configured to receive input indicating an operating mode; a power switch configured to supply power to a controller upon receiving a signal from the user input device; and a controller configured to control power supplied to one or more subsystems of the aircraft. The user input device is configured to: not send the signal to the power switch when the received input indicates a first operating mode; send the signal to the power switch when the received input indicates a second operating mode; send the signal to the power switch when the received input indicates a third operating mode; and not send the signal to the power switch when the received input indicates a fourth operating mode. Attached Figure Description

[0008] Figure 1A An example eVTOL aircraft consistent with embodiments of this disclosure is shown.

[0009] Figure 1B Another example eVTOL aircraft consistent with embodiments of this disclosure is shown.

[0010] Figure 1C , Figure 1D , Figure 1E , Figure 1F , Figure 1G and Figure 1H An exemplary top plan view of an aircraft consistent with the disclosed embodiments is shown.

[0011] Figure 1I An example eVTOL aircraft and associated subsystems consistent with embodiments of this disclosure are shown.

[0012] Figure 1J , Figure 1K and Figure 1L Exemplary connections to subsystems consistent with embodiments of this disclosure are shown.

[0013] Figure 2A An exemplary mode of an aircraft consistent with embodiments of this disclosure is shown.

[0014] Figure 2B , Figure 2C , Figure 2D and Figure 2E Exemplary electrical connections associated with aircraft modes are shown, consistent with embodiments of this disclosure.

[0015] Figure 3A A diagram illustrating an exemplary aircraft circuit system consistent with embodiments of this disclosure is shown.

[0016] Figure 3B A diagram illustrating an exemplary aircraft circuit system consistent with embodiments of this disclosure is shown.

[0017] Figure 3C A diagram showing an exemplary mode switch consistent with embodiments of this disclosure is provided.

[0018] Figure 3D Another diagram shows an exemplary mode switch consistent with embodiments of this disclosure.

[0019] Figure 4 A flowchart illustrating an exemplary process for switching aircraft modes, consistent with embodiments of this disclosure, is shown.

[0020] Figure 5 A flowchart illustrating an exemplary process for switching aircraft states, consistent with embodiments of this disclosure, is shown. Detailed Implementation

[0021] This disclosure presents systems, components, and technologies primarily for use in aircraft. An aircraft can be a manned aircraft, an unmanned aircraft (e.g., a UAV), a drone, a helicopter, and / or an airplane. An aircraft comprises a fuselage and one or more components configured to allow flight (e.g., wings, tail, propeller). An aircraft may contain any configuration that requires power to one or more subsystems of the aircraft.

[0022] The disclosed embodiments provide new and improved configurations for some aircraft components not observed in conventional aircraft, and / or identify design standards for components that differ from those in conventional aircraft. These alternative configurations and design standards address the shortcomings and challenges of conventional components, resulting in the embodiments disclosed herein of various configurations and designs for components of propulsion-driven aircraft (e.g., electric or hybrid electric aircraft).

[0023] Embodiments may include an electric propulsion system comprising an electric engine connected to an onboard electric power source. The power source may include devices capable of storing energy, such as batteries or capacitors, and may optionally include one or more systems for utilizing or generating electricity, such as a fuel-powered generator or a solar panel array. In some embodiments, the aircraft may include a hybrid-electric aircraft that uses at least one of an electric-based energy source or a fuel-based energy source to power the distributed propulsion system. In some embodiments, the aircraft may be powered by one or more batteries, an internal combustion engine (ICE), a generator, a turbine engine, or a ducted fan.

[0024] In some embodiments, the electric propulsion system as described herein can generate thrust by supplying high-voltage (HV) electrical power to an electric engine, which in turn converts the HV electrical power into mechanical shaft power to rotate the propeller. An aircraft as described herein may have multiple electric engines. The magnitude of the thrust generated by each electric engine can be controlled by torque commands sent from the flight control system (FCS) to each electric engine via a digital communication interface.

[0025] Embodiments of this disclosure implement improved system redundancy in the event of a failure, minimizing any individual point of failure in the aircraft propulsion system. Some of the disclosed embodiments also provide new and improved methods to meet aviation and transport laws and regulations.

[0026] Aircraft need to be lightweight enough to meet their performance objectives. Additional weight on an aircraft can affect its ability to generate sufficient lift (e.g., through static and / or powered lift elements) for safe flight. Furthermore, additional weight can negatively impact an aircraft's maneuverability, speed, rate of climb, and range. Of particular importance is that electric aircraft avoid additional weight. Unlike conventional aircraft, electric aircraft cannot store additional fuel. Instead, the energy supply of an electric aircraft is limited by the energy capacity of its battery pack.

[0027] Battery packs are one of the major contributors to the weight of electric aircraft. Electric aircraft can contain multiple high-voltage battery packs to power one or more electric engines and / or tilt actuators. For example, an electric aircraft can contain four, six, eight, ten, or twelve, or any number of battery packs, to power the aircraft's electric engines and / or tilt actuators. One or more additional battery packs (e.g., low-voltage battery packs) can power low-voltage systems such as flight control systems, charging control equipment, environmental conditioning systems, and / or other auxiliary payloads. However, these additional battery packs add weight and occupy space within the aircraft.

[0028] Embodiments of this disclosure help solve this and other problems by utilizing one or more high-voltage battery packs to power a low-voltage power distribution system, which also powers the electric motor and / or tilt actuator. In some embodiments, one or more DC / DC converters are configured to reduce the high-voltage power to feed it into the low-voltage power distribution system. Therefore, the number of battery packs can be reduced and / or the need for low-voltage battery packs can be eliminated.

[0029] Furthermore, embodiments of this disclosure provide arrangements of components and circuitry that allow for individual control of power to different loads (e.g., low-voltage systems, electric motors, etc.), thereby improving energy efficiency. Embodiments of this disclosure also provide redundancy and fault tolerance by balancing electrical isolation between backup power supplies and critical components.

[0030] Furthermore, embodiments of this disclosure allow pilots to control the aircraft in different modes with certain subsystems enabled or disabled. For example, a pilot might want to control the aircraft in different modes, such as "off" mode, "maintenance mode," "ground mode," and / or "flight" mode, where different subsystems are enabled (e.g., for each mode). However, pilots may find it difficult to turn different subsystems on and off. This disclosure addresses this and other problems by providing a mode switch that automatically enables and / or disables different subsystems based on the selected mode. Additionally, within each mode, the aircraft can be automatically controlled to different states based on detected events.

[0031] Furthermore, embodiments of this disclosure help limit the operating hours of critical aircraft systems, such as the Flight Control System (FCS). Limiting FCS operating hours can help maintain FCS integrity and / or regulatory compliance. This disclosure addresses this and other issues by deactivating the FCS in certain modes and / or states and relying on one or more other controllers (e.g., a charging control unit) to communicate with various subsystems. For example, in some embodiments, the FCS can be deactivated in "shutdown" and "maintenance" modes unless a maintenance state is selected and / or the ground unit supplies power to the low-voltage power distribution system. Limiting FCS usage can also reduce unnecessary power consumption.

[0032] Furthermore, embodiments of this disclosure help limit the operating hours of other low-voltage components (such as one or more low-voltage system control units and / or one or more low-voltage power distribution units) to maintain their integrity. This disclosure addresses this and other problems by disabling other low-voltage components in certain modes and / or states. For example, in some embodiments, low-voltage components can be disabled in "shutdown" and "maintenance" modes unless a maintenance state is selected and / or the ground unit supplies power to the low-voltage distribution system. Limiting the use of low-voltage components also reduces unnecessary power consumption.

[0033] This disclosure further allows charging in certain modes where the FCS and / or other low-voltage components are deactivated. For example, in some embodiments, the FCS and / or other low-voltage components can be deactivated in "off" and "maintenance" modes, and charging can be permitted in these modes. Therefore, aircraft charging (e.g., overnight) can be performed without increasing the additional operating hours on the FCS and / or other low-voltage components.

[0034] Examples of the embodiments are now illustrated in detail with reference to the accompanying drawings. The following description refers to the accompanying drawings, in which, unless otherwise indicated, the same numerals in different figures represent the same or similar elements. The embodiments set forth in the following description of the exemplary embodiments do not represent all embodiments consistent with this disclosure. Instead, the embodiments are merely examples of apparatuses and methods consistent with aspects related to the subject matter recited in the appended claims.

[0035] Figure 1A-Figure 1B VTOL aircraft 100, consistent with embodiments of this disclosure, is shown in both a cruise configuration and a vertical takeoff, landing, and hovering configuration (also referred to herein as a "lift" configuration). Aircraft 100 may include a fuselage 108, wings 109 mounted to the fuselage 108, a tail 107, and one or more rear stabilizers 106 mounted to the rear of the tail 107 or fuselage 108. A plurality of lift propellers 112 may be mounted to the wings 109 and configured to provide lift for vertical takeoff, landing, and hovering. A plurality of tilt propellers 114 may be mounted to the wings 109 and may tilt between the cruise and lift configurations, in which the tilt propellers provide forward thrust to the aircraft 100 for horizontal flight, such as... Figure 1A As shown, in the lift configuration, the plurality of tilting propellers provide a portion of the lift required for vertical takeoff, landing, and hovering, such as Figure 1B As shown. As used herein, lift configuration can refer to the orientation of a tilting propeller whose thrust primarily provides lift to the aircraft. Cruise configuration can refer to the orientation of a tilting propeller whose thrust primarily provides forward thrust to the aircraft. Alternatively, cruise configuration can refer to a configuration in which the lift propeller is retracted.

[0036] In some embodiments, the lift propeller 112 can be configured to provide lift only, with all propulsion provided by the tilt propeller. Thus, the lift propeller 112 can be in a fixed position and can generate thrust only during takeoff, landing, and hovering. Simultaneously, the tilt propellers 114 can tilt into a lift configuration in which their thrust is directed downwards to provide additional lift.

[0037] For forward flight, the tilt propellers 114 can tilt from their lift configuration to their cruise configuration. In other words, the pitch and tilt angle of the tilt propellers 114 can vary from a tilt propeller thrust-directed downward direction (to provide lift during vertical takeoff, landing, and hovering) to a tilt propeller thrust-directed rearward direction (to provide forward thrust to the aircraft 100). The tilt propellers can tilt about an axis that can be perpendicular to the forward direction of the aircraft 100. When the aircraft 100 is in fully forward flight during cruise configuration, lift can be provided entirely by the wing 109. Meanwhile, the lift propellers 112 can be shut down. The blades 121 of the lift propellers 112 can be locked in a low-drag position for aircraft cruise. In some embodiments, each lift propeller 112 may have two blades 121, which can be locked for cruise in a minimum drag position, where one blade is directly in front of the other, such as... Figure 1A As shown. In some embodiments, the lift propeller 112 has more than two blades. In some embodiments, the tilt propeller 114 includes more blades 118 than the lift propeller 112. For example, as Figure 1A-Figure 1B As shown, each of the lift propellers 112 may include, for example, two blades, and each of the tilt propellers 114 may include, for example, five blades. In some embodiments, the tilt propellers 114 may have, for example, two to five blades.

[0038] In some embodiments, the aircraft may include only one wing 104 on each side of the fuselage 108 (or a single wing extending across the entire aircraft), and at least a portion of the lift propeller 112 may be located behind the wing 109, and at least a portion of the jib propeller 114 may be located forward of the wing 109. In some embodiments, all lift propellers 112 may be located behind the wing 109, and all jib propellers 114 may be located forward of the wing 109. According to some embodiments, all lift propellers 112 and jib propellers 114 may be mounted on the wing, i.e., no lift propeller or jib propeller may be mounted on the fuselage. In some embodiments, all lift propellers 112 may be located behind the wing 109, and all jib propellers 114 may be located forward of the wing 109. According to some embodiments, all lift propellers 112 and jib propellers 114 may be located inside the wingtip 109.

[0039] In some embodiments, the lift propeller 112 and the tilt propeller 114 can be mounted to the wing 109 via a boom 122. The boom 122 can be mounted below the wing 109, on the top of the wing, and / or integrated into the wing profile. In some embodiments, each boom 122 can mount one lift propeller 112 and one tilt propeller 114. The lift propeller 112 can be mounted at the rear end of the boom 122, and the tilt propeller 114 can be mounted at the front end of the boom 122. In some embodiments, the lift propeller 112 can be mounted in a fixed position on the boom 122. In some embodiments, the tilt propeller 114 can be mounted to the front end of the boom 122 via a hinge. The tilt propeller 114 can be mounted to the boom 122 such that, in a cruise configuration, the tilt propeller 114 is aligned with the body of the boom 122, thereby forming a continuous extension at the front end of the boom 122 that minimizes drag during forward flight.

[0040] In some embodiments, aircraft 100 may include, for example, a wing on each side of fuselage 108 or a single wing extending across the aircraft. According to some embodiments, at least one wing 104 is a high wing mounted to the upper side of fuselage 108. According to some embodiments, the wing includes control surfaces such as flaps, ailerons, or flaperones. Further discussion of VTOL aircraft can be found in U.S. Patent Publication No. 2021 / 0362849, the entire contents of which are incorporated herein by reference.

[0041] Figure 1C , Figure 1D , Figure 1E , Figure 1F , Figure 1G and Figure 1H An exemplary top plan view of an aircraft consistent with the disclosed embodiments is shown. There may be numerous design considerations (cost, weight, size, performance capabilities, etc.) that could affect the number and / or combination of tilting (e.g., tiltable) propellers and lift (e.g., non-tilt) propellers in an aircraft. The number and orientation of propulsion units can affect the number of battery packs and the connections between them (e.g., to achieve controllability and / or stability in the event of electrical failure).

[0042] Figure 1C An arrangement of the electric propulsion unit consistent with embodiments of this disclosure is shown. Reference Figure 1CThe aircraft shown in the figure may be a top-down plan view of an exemplary aircraft. The aircraft may include twelve electric propulsion systems distributed across the aircraft. In some embodiments, the distribution of electric propulsion systems may include six forward electric propulsion systems (165, 166, 167, 168, 169, and 170) and six tail electric propulsion systems (171, 172, 173, 174, 175, and 176). In some embodiments, the six forward electric propulsion systems may be operatively connected to a tilting propeller, and the six tail electric propulsion systems may be operatively connected to a lift propeller. In other embodiments, the six forward electric propulsion systems and multiple tail electric propulsion systems may be operatively connected to the tilting propeller, and the remaining tail electric propulsion systems may be operatively connected to the lift propeller. In other embodiments, all forward and tail electric propulsion systems may be operatively coupled to the tilting propeller.

[0043] Figure 1D An alternative arrangement of the electric propulsion unit consistent with embodiments of this disclosure is shown. Reference Figure 1D The aircraft shown in the figure may be a top-down plan view of an exemplary aircraft. The aircraft may include eight electric propulsion systems distributed across the aircraft. In some embodiments, the distribution of electric propulsion systems may include four forward electric propulsion systems (177, 178, 179, and 180) and four tail electric propulsion systems (181, 182, 183, and 184). In some embodiments, the four forward electric propulsion systems may be operatively connected to a tilting propeller, and the four tail electric propulsion systems may be operatively connected to a lift propeller. In other embodiments, the four forward electric propulsion systems and multiple tail electric propulsion systems may be operatively connected to a tilting propeller, and the remaining tail electric propulsion systems may be operatively connected to a lift propeller. In other embodiments, all forward and tail electric propulsion systems may be operatively coupled to a tilting propeller.

[0044] Figure 1E An alternative arrangement of the electric propulsion unit consistent with embodiments of this disclosure is shown. Reference Figure 1EThe aircraft may be a top-down plan view of an exemplary aircraft. In some embodiments, the aircraft may include ducted fans operatively connected to an electric propulsion system. In some embodiments, the aircraft may include a set of ducted fans on each wing of the aircraft, and the set of ducted fans may be connected together to tilt (e.g., between a lift configuration and a forward thrust configuration). In some embodiments, the aircraft includes a left and right forewing and a left and right aft wing. In some embodiments, each wing of the aircraft includes a set of connected ducted fans. In some embodiments, each set of connected ducted fans is tiltable (e.g., between lift and forward thrust), while in other embodiments, only one set of fans on one or more forewings is tiltable.

[0045] Figure 1F An alternative arrangement of the electric propulsion unit consistent with embodiments of this disclosure is shown. Reference Figure 1F The aircraft shown in the figure may be a top plan view of an exemplary aircraft. The aircraft may include six electric propulsion systems distributed across the aircraft. In some embodiments, the distribution of the electric propulsion systems may include a first set of four electric propulsion systems 185, 186, 187, and 188 coplanar in a first plane and a second set of two electric propulsion systems 189 and 190 coplanar in a second plane. In some embodiments, the first set of electric propulsion systems 185, 186, 187, and 188 may be operatively connected to a tilting propeller, and the second set of electric propulsion systems 189 and 190 may be operatively connected to a lift propeller. In other embodiments, the first set of electric propulsion systems 185, 186, 187, and 188 and the second set of tail-end electric propulsion systems 189 and 190 may all be operatively connected to the tilting propeller.

[0046] Figure 1G An alternative arrangement of the electric propulsion unit consistent with embodiments of this disclosure is shown. Reference Figure 1G The aircraft shown in the figure may be a top-down plan view of an exemplary aircraft. The aircraft may include four electric propulsion systems distributed across the aircraft. In some embodiments, the distribution of electric propulsion systems may include four coplanar electric propulsion systems 191, 192, 193, and 194. In some embodiments, all electric propulsion systems may be operatively connected to tilting propellers.

[0047] Figure 1H An alternative arrangement of the electric propulsion unit consistent with embodiments of this disclosure is shown. Reference Figure 1HThe aircraft shown in the figure may be a top plan view of an exemplary aircraft (e.g., a VTOL aircraft). The aircraft may include six electric propulsion systems spanning the aircraft's distribution. For example, in some embodiments, the aircraft may include four forward electric propulsion systems 195, 196, 197, and 198 operably connected to a tilting propeller and two tail electric propulsion systems 199 and 200 operably connected to a lift propeller. In some embodiments, the aircraft may include ten electric propulsion systems spanning the aircraft's distribution. For example, in some embodiments, the aircraft may include six forward electric propulsion systems operably connected to a tilting propeller and four tail electric propulsion systems operably connected to a lift propeller. In some embodiments, some or all of the tail electric propulsion systems may be operably connected to the tilting propeller.

[0048] like Figure 1H As shown, in some embodiments, the aircraft may have a flying wing configuration, such as a tailless fixed-wing aircraft without a fixed fuselage. In some embodiments, the aircraft may have a flying wing configuration in which the fuselage is integrated into the wing. In some embodiments, when the tiltrotor is operated in a lift configuration, the tiltrotor can rotate in a plane above the aircraft fuselage.

[0049] Figure 1I An example eVTOL aircraft and associated subsystems consistent with embodiments of this disclosure are shown. Subsystem 101 is a high-voltage circuitry system and associated devices that power the electric engine and tilt actuator. Subsystem 102 is a high-voltage circuitry system and associated devices connected to a DC / DC converter that reduces power to feed a low-voltage power distribution system. In some embodiments, each battery pack in the aircraft may provide power to both subsystem 101 and subsystem 102. In some embodiments, the arrangement of switching devices provides individual control over the power flow to subsystem 101 and subsystem 102.

[0050] Subsystem 103 is a low-voltage power distribution system and associated devices. As described above, the low-voltage power distribution system can be fed from subsystem 102 via a DC / DC converter. For example, one or more primary and / or alternative DC / DC converters can perform voltage conversion. In some embodiments, an alternative converter can provide low-voltage power during aircraft power-on and / or in the event of a failure of one or more primary DC / DC converters. Furthermore, subsystem 103 can be configured to receive power via a ground power unit. In some embodiments, only one of subsystems 103 or 104 can be configured to receive power from a ground unit at any given time. In other embodiments, both subsystems can be configured to receive power from one or more ground units simultaneously.

[0051] The low-voltage power distribution system (subsystem 103) may include a combination of power distribution boxes (PDBs) to manage and distribute low-voltage power to different aircraft components. In some embodiments, the low-voltage power distribution system can power all low-voltage aircraft components. For example, the low-voltage power distribution system can power avionics, flight control computers, aircraft flight control surfaces, motor controllers, battery management systems for battery packs, environmental control systems, sensors, medical devices, and / or any other systems on the aircraft that require low-voltage power. As used herein, “high voltage” means a voltage greater than 110 volts. As used herein, “low-voltage DC” means 33 VDC or lower. In some embodiments, low voltage means 28 VDC.

[0052] Subsystem 104 is a high-voltage charging bus and associated devices that connect charging ports to one or more battery packs to allow the battery packs to be charged. In some embodiments, the high-voltage charging bus may be electrically isolated from other high-voltage power distribution. For example, the charging bus may be electrically isolated from the high-voltage power supplied to the electric motor and actuators (subsystem 101) and the high-voltage power supplied to the converter (subsystem 102) for low-voltage power distribution.

[0053] Subsystem 105 is a charging control unit (CCU). The CCU can control the battery pack during the charging process. For example, in some embodiments, the CCU can receive status updates from the battery pack and provide commands to the battery pack to control their charging levels by opening and closing the battery pack charging contactors. Furthermore, in some embodiments, the CCU can send commands to one or more ground units to transmit the battery pack's charging and / or cooling requirements.

[0054] Furthermore, in some embodiments, the CCU can receive the selected mode and communicate with one or more components of the aircraft. For example, in some embodiments, the CCU can transmit the selected mode to one or more battery packs and control the battery packs as needed. In some embodiments, the communication line between the CCU and the battery pack can be used for both mode switching control and charging control. Therefore, the amount of wiring in the aircraft can be reduced. In other embodiments, a controller other than the CCU can receive the selected mode and perform communication with components of the aircraft. For example, in some embodiments, the aircraft may include a central battery management unit and / or a central battery management system, which controls the operation of all aircraft battery packs (e.g., sending contactor commands, monitoring battery pack status, etc.). The central battery management unit and / or battery management system can receive the selected mode and perform the communication detailed above. Additionally, a separate controller (e.g., a central battery management unit or BMS) can control the charging operation.

[0055] Figure 1J , Figure 1K and Figure 1L Exemplary connections to subsystems consistent with embodiments of this disclosure are shown. Figure 1J This illustrates a battery pack configured to power the electric propulsion unit and provide backup power to other battery packs. The configuration and control of the electric engine and propeller can be matched... Figure 1A-Figure 1B The aircraft configuration and control described herein. The aircraft may include a High Voltage Power Supply (HVPS) system to supply high voltage (HV) electrical power. The HVPS system is a power source on the aircraft and is configured to distribute stored electrical energy to other systems on the aircraft, including an Electric Propulsion System (EPS) for converting electrical power into mechanical rotating shaft power to generate thrust. As shown, the aircraft's HVPS system may include six battery packs 120 (numbered 1-6 from left to right) mounted in battery trays in the aircraft's wings. The battery packs can power one or more electric engines 110. Although six battery packs are shown, the aircraft 100 may have any number of battery packs.

[0056] In some embodiments, a single battery pack 120 may be electrically connected to and power a plurality of electric motors. For example, in some embodiments, the battery pack 120 may power an electric motor 110 on either side of a longitudinal axis. In some embodiments, the battery pack 120 may power an electric motor on either side of a horizontal axis. In some embodiments, as shown, the battery pack 120 may power two diagonally opposite electric motors 110.

[0057] Furthermore, the HVPS system may include cross-connect lines 130 (e.g., electrical connections between battery packs) having at least one fuse allowing pairing of two or more battery packs 120. Power for the electric engines 110 can be shared among paired battery packs 120 via the cross-connect lines 130. Therefore, multiple battery packs 120 can simultaneously power multiple electric engines 110. This arrangement provides redundancy and avoids a single point of failure, as each paired battery pack 120 can serve as a backup for other battery packs. In the event of a battery pack failure, one or more connected battery packs 120 can continue to power the connected electric engines 110 of the failed battery pack. In some embodiments, as shown, two battery packs 120 may be paired together to power a total of four electric engines 110. However, the aircraft may include different combinations of electric engine and battery pack pairings.

[0058] Alternative battery pack configurations: The above configuration is provided as an example, but different numbers and configurations of battery packs, electric motors, battery pack-to-electric motor connections, and battery pack cross-connections can be used. In some embodiments, each battery pack can power an individual electric motor. For example, an aircraft can have four, six, eight, ten, twelve, or any number of electric motors, and the number of battery packs can match the number of electric motors. In some embodiments, each battery pack can power only one electric motor and can be electrically isolated from all other battery packs. In some embodiments, each battery pack can power one or more partial motors, and each electric motor can contain two or more partial motors. Therefore, each electric motor can have a backup power source, but the battery packs can still be electrically isolated.

[0059] In some embodiments, each battery pack can power multiple electric engines. As described above, the battery packs can power multiple sets of electric engines symmetrically arranged along one or more axes of symmetry. In some embodiments, the battery packs can power electric engines symmetrically arranged along the longitudinal axis, lateral axis, or both of the aircraft's axis. For example, as described above, in some embodiments, different battery packs can power diagonally symmetrical electric engines.

[0060] In some embodiments, the battery pack can power more than two electric motors. In some embodiments, the battery pack can power two or more sets of diagonally symmetrical electric motors. In some embodiments, a set of electric motors powered by the battery pack may include a pair of electric motors that are diagonally symmetrical on the inner side and a pair of electric motors that are diagonally symmetrical on the outer side. In some embodiments, the battery pack can power four or more electric motors that are symmetrically arranged along a longitudinal axis of symmetry.

[0061] In some embodiments, some or all of the battery packs are interconnected. As described above, cross-connection lines can allow each battery pack to act as a backup power source for another battery pack. For example, in some embodiments, a first battery pack can directly power a first number of electric motors, and a second battery pack can directly power a second number of electric motors. The first and second battery packs can be cross-connected together to form a battery pack unit. Thus, each battery pack in the unit can be used as a backup for another battery pack. In the event of a battery pack failure in the unit, the failed battery pack can be disconnected, and the electric motors will be powered by one or more non-failed battery packs in the unit. The battery packs in the battery pack unit can be electrically isolated from other battery pack units.

[0062] In some embodiments, a battery pack unit may include three battery packs, each powering multiple electric motors. For example, in some embodiments, each battery pack may power two diagonally symmetrical electric motors. Therefore, each battery pack unit can power a total of six electric motors, with each electric motor having two backup battery packs. In some embodiments, each battery pack unit may power four electric motors, comprising two sets of diagonally symmetrical electric motors. Therefore, each battery pack unit can power a total of twelve electric motors, with each electric motor having two backup battery packs.

[0063] In some embodiments, the battery pack unit may include four battery packs, each powering multiple electric motors. For example, in some embodiments, each battery pack may power two diagonally symmetrical electric motors. Therefore, each battery pack unit can power a total of eight electric motors, with each electric motor having three backup battery packs. In other embodiments, each battery pack unit may power four electric motors, comprising two sets of diagonally symmetrical electric motors. Therefore, each battery pack unit can power sixteen electric motors, with each electric motor having three backup battery packs.

[0064] In some embodiments, all battery packs are cross-connected to their neighbors to form a ring power supply for the battery packs. In some embodiments, all battery packs are connected to a common bus. In some embodiments, the common bus may form a ring power supply to provide additional connectivity redundancy; while in other embodiments, the common bus may not form a ring power supply.

[0065] In some embodiments, the electric motor includes a single motor powered by one or more battery packs. In some embodiments, each electric motor may include two or more partial motors, and the battery packs may power the partial motors. In some embodiments, any electric motor power supply configuration described above may include powering one or more partial motors that power one or more battery packs.

[0066] Different configurations can be selected, including combinations of battery packs, electric engines, battery pack-to-electric engine connections, and battery pack cross-connections, to optimally balance aircraft power requirements, system redundancy, and fault tolerance. For example, different configurations can be used to accommodate the above. Figure 1C-Figure 1H Details of the aircraft shown.

[0067] Figure 1KThe battery packs can also power the tilting propeller system of the electric propulsion unit. For example, the battery packs can power linear actuators and / or rotary actuators to change the orientation of the propulsion system (e.g., the tilt angle of one or more tilting propellers) during operation (represented by T1-T6). In some embodiments, the rotary actuator may include a motor, an inverter, and a gearbox. In some embodiments, as shown, each battery pack powers a tilting propeller system corresponding to an electric engine powered by the battery pack. As described above, aircraft may include varying numbers and / or combinations of battery packs and propulsion system arrangements (with corresponding tilting propeller systems). In some embodiments, reference... Figure 1J Each of the described arrangements may further include a battery pack configured to supply power to the tilt actuator in addition to powering the tilt-controlled electric propulsion unit (e.g., an electric engine).

[0068] Figure 1L The battery pack may also have power to operate to one or more DC / DC converters to supply low-voltage power to low-voltage components of the aircraft. Furthermore, one or more battery packs may power environmental conditioning equipment, such as compressors, fans, or other equipment requiring high-voltage power. In some embodiments, the low-voltage system is powered without a separate low-voltage battery, while in other embodiments, one or more low-voltage components may be powered by a separate low-voltage battery.

[0069] Figure 2A The diagram illustrates an aircraft mode consistent with embodiments of this disclosure. As further discussed herein, consistent with the disclosed embodiments, at least one aircraft component (e.g., at least one controller, at least one processor, at least one battery management unit, at least one mode switch, or any combination thereof) connected to one or more electrical systems of the aircraft can control the power distribution to one or more components and / or subsystems. In some embodiments, the mode switch (or other component) can distribute power based on (e.g., in response to, limited to) a selected mode (e.g., selected by a user at an interface such as a graphical user interface or a physically moving part, consistent with the disclosed embodiments). Additionally, consistent with the disclosed embodiments, the mode switch can also distribute power based on the state of one of a plurality of possible states present in a particular mode.

[0070] In some embodiments, the mode switch can be configured to distribute power in different configurations (e.g., each configuration associated with a corresponding mode) (e.g., flowing or not flowing). In some embodiments, a selectable mode can be associated with connecting (or maintaining a power connection) power to a set of components (e.g., at least one subsystem, at least one low-voltage subsystem, at least one high-voltage subsystem, at least one charging subsystem, at least one bus, or any combination thereof) while simultaneously disconnecting (or maintaining a power disconnect) power to another set of components. In some embodiments, another selectable mode can be associated with connecting power to multiple subsystems (e.g., all systems required for aircraft flight). In some embodiments, the mode switch can be configured to implement one or more of the modes described herein.

[0071] In shutdown mode 201, subsystem 101 is deactivated, and no high-voltage power will run to the electric motor and tilt actuator. Subsystem 102 is deactivated, and no high-voltage power will flow from the battery pack to the DC / DC converter to feed power to the low-voltage power distribution system. Subsystem 103 is deactivated, and no low-voltage power will be distributed through the low-voltage power distribution system. Subsystem 104 is deactivated, and no charge will be applied to the battery pack. Subsystem 105 is deactivated, and the CCU will remain unpowered. Aircraft maintenance testing is unavailable in shutdown mode 201.

[0072] In maintenance mode 202, subsystem 101 is deactivated, and no high-voltage power will run to the electric motor and tilt actuator. Subsystem 102 is deactivated, and no high-voltage power will flow from the battery pack to the DC / DC converter to feed power to the low-voltage distribution system. Subsystem 103 is activated, but low-voltage power can be obtained solely through the ground power unit instead of through subsystem 102. Subsystem 104 is activated, and the battery pack can be charged. Subsystem 105 is activated, and the CCU will be powered.

[0073] Furthermore, in maintenance mode 202, service mode maintenance is available when the low-voltage power distribution system (i.e., subsystem 103) is powered via a ground unit. For example, aircraft components such as the system control unit (SCU) used to control low-voltage power distribution can receive signals indicating maintenance mode 202 and can enable service mode maintenance. For example, the SCU and / or one or more other aircraft components can implement software loading. In maintenance mode 202, software can be loaded onto one or more battery management units. Additionally, the SCU and / or one or more other aircraft components can enable the testing of low-voltage equipment. Low-voltage equipment testing can include testing avionics, flight control computers, aircraft flight control surfaces, motor controllers, pitch actuators, tilt actuators, battery management systems for battery packs, environmental control systems, sensors, and / or any other systems on the aircraft that require low-voltage power.

[0074] In ground mode 203, when performing maintenance tests, subsystem 101 is activated, containing high-voltage power supplied to the electric motor and tilt actuators. Subsystem 102 is activated, and power flows from the battery pack to a DC / DC converter to feed power to the low-voltage power distribution system. Subsystem 103 is activated, and the low-voltage power is distributed through the low-voltage power distribution system. The low-voltage power can be obtained via a DC / DC converter (e.g., via subsystem 102) or via one or more ground units. Subsystem 104 is activated, and the battery pack can be charged. Subsystem 105 is activated, and the CCU is powered.

[0075] Furthermore, ground mode maintenance is available in ground mode 203. For example, an aircraft component such as a system control unit (SCU) for controlling low-voltage power distribution can receive a signal instructing ground mode 203 and can enable ground mode maintenance. Ground mode maintenance may involve loading software into different aircraft systems. In some embodiments, ground mode maintenance may prevent software loading into the battery management unit (BMU). The SCU and / or one or more aircraft components may allow testing of low-voltage equipment. Low-voltage equipment testing may include testing avionics, flight control computers, aircraft flight control surfaces, motor controllers, pitch actuators, tilt actuators, battery management systems for battery packs, environmental control systems, sensors, and / or any other systems on the aircraft that require low-voltage power.

[0076] In some embodiments, the maintenance mode can be selected via a user input device. For example, the maintenance mode can be selected via a physical switch, button, joystick, and / or display element. The user input device can be installed on the aircraft fuselage and / or in the pilot's cockpit. In some embodiments, the maintenance mode can be selected via a mode switch. As further described below, the selection of the maintenance mode can be transmitted to the CCU.

[0077] Furthermore, the FCS and / or other components can receive signals instructing ground-mode maintenance tests and can allow testing of high-voltage equipment. The FCS and / or other components facilitate testing of electric engines and associated electric propulsion units, such as rotors, propellers, and / or propellers. The FCS and / or other components facilitate testing of tilting propeller systems, including linear actuators and / or rotary actuators.

[0078] In flight mode 204, subsystem 101 is activated, and high-voltage power will flow to the electric motors and tilt actuators. Subsystem 102 is activated, and power will flow from the battery pack to the DC / DC converter to feed the low-voltage power distribution system. Subsystem 103 is activated, and low-voltage power will be obtained through the DC / DC converter and associated battery pack. Subsystem 104 is deactivated, and no charge will be applied to the battery pack. Subsystem 105, the CCU, is deactivated, and the CCU will remain unpowered. Aircraft maintenance testing is unavailable in flight mode 204.

[0079] In some embodiments, the aircraft may be prohibited from flying in shutdown mode, maintenance mode, and / or ground mode. For example, the flight control system may prevent the aircraft from being controlled in a manner that allows takeoff.

[0080] While four modes are provided as examples, an aircraft can contain a different number of modes. For example, an aircraft can contain two modes, including shutdown mode and flight mode, or maintenance mode and flight mode. For example, an aircraft can contain three modes, including shutdown mode, ground mode, and flight mode. For example, an aircraft can contain five modes, in addition to... Figure 2A In addition to the modes shown, charging modes are also included. For example, an aircraft may include additional modes based on the intended flight type. For instance, one mode may provide the high-voltage power required for routine takeoff and landing missions, while another mode may provide power for powered lift missions (e.g., providing power to a lift engine). In some embodiments, when employing different numbers of modes, it may be necessary to check different conditions before allowing a mode switch. For example, an aircraft with maintenance and flight modes may check the connectivity of the DC / DC converter before allowing a switch to flight mode. An aircraft with different modes depending on the flight type may check the connectivity of the electric propulsion unit and / or tilt actuator for that flight type.

[0081] As shown, power supplied to different subsystems can be controlled based on mode selection. For example, in some embodiments, high voltage is controlled differently in at least three operating modes (e.g., different HV circuits connected to the power source). In some embodiments, low voltage is controlled differently in at least two operating modes (e.g., standby and ground) (e.g., different LV circuits connected to the power source). In some embodiments, low voltage is controlled in the same way in at least two operating modes (e.g., ground and flight) (e.g., substantially the same LV distribution lines connected to the power source).

[0082] Figure 2B , Figure 2C , Figure 2D and Figure 2E The diagram illustrates the electrical configuration associated with an aircraft mode, consistent with embodiments of this disclosure. In some embodiments, mode selection is transmitted to the Battery Management Unit (BMU) of each battery pack, which is configured to respond by opening and / or closing one or more switching devices. The BMU may include one or more computers, processors, microprocessors, controllers, and / or associated circuitry. One or more mode selections may be transmitted to the BMU from a central charging control unit, flight control system, and / or one or more different components of the aircraft. Furthermore, in some embodiments, the BMU, FCS, and / or another component may confirm that one or more conditions are met before switching to the selected mode. After confirmation, the BMU may control one or more switching devices to enable and / or disable certain subsystems. One or more switching devices K1-K7 may include one or more contactors, one or more relays, one or more transistors, one or more controllers, and / or any other device capable of switching power on and off. In some embodiments, K1-K7 are all contactors. As detailed below, switching devices K2-K7 may control whether one or more battery packs receive charge from subsystem 104, which includes a charging bus. Switching devices K1-K5 can control whether high-voltage power flows to the electric motor, tilt actuator, and / or DC / DC converter. Therefore, the switching devices can enable and / or disable subsystem 101, including the high-voltage power supplied to the electric motor and tilt actuator. Furthermore, the contactors can enable and / or disable subsystem 102, including the high-voltage power supplied to the DC / DC converter used for low-voltage power distribution.

[0083] Furthermore, as shown, each battery pack may include a high-voltage junction box (HVJB) electrically connected to the HV loads to provide high-voltage power. Specifically, energy storage elements BT1 (e.g., battery cells connected in parallel and series) may be used to provide high-voltage power. The energy storage elements BT1 are connected to each HV load via one or more pre-charge resistors (e.g., resistors R1 and R7) or one or more current-sensing resistors (e.g., resistors R2-R6), switching devices K1-K5 (e.g., HV contactors), and a combination of active and passive fuses (e.g., F1-F7) to prevent various fault conditions (e.g., overcurrent, short circuit, etc.). In some embodiments, different configurations of active and passive fuses can prevent fault conditions. For example, a single active fuse may be included in the cross-connection lines between battery packs.

[0084] exist Figure 2B The diagram shows the "off" mode, where no high-voltage power is supplied to aircraft components. The BMU can command switches K1-K7 to disconnect (e.g., activate). Furthermore, the BMU is powered by the BMU DC / DC converter, allowing it to periodically wake up and monitor individual battery cells (V, T, I), estimate HV isolation, perform balancing, and / or detect latch-up faults. This is further described below. Figure 2B It also demonstrates a "maintenance" mode, in which the CCU (not shown) is powered via an internal DC / DC converter. Figure 2C The "Maintenance" mode indicates that charging can be enabled.

[0085] exist Figure 2D The diagram illustrates the "ground" mode, where high-voltage power is supplied to the DC / DC converter and ECS system, allowing power to be supplied to low-voltage systems and environmental control equipment. The BMU can command switching device K3 to close (e.g., deactivate) to allow pre-charging of the circuitry and limit inrush current to the DC / DC converter and ECS equipment circuitry. Once the BMU determines that pre-charging is complete (e.g., a set time has elapsed or the voltage or stored charge has risen to a threshold level), the BMU will disconnect K3 and close K1 and K2, allowing high-voltage power from the battery pack to feed power to the DC / DC converter and / or ECS equipment. In some embodiments, by closing K1 and K2, power is also supplied to cross-connect lines (e.g., cross-connect line 130) to one or more connected battery packs, providing backup power. Furthermore, the BMU can estimate high-voltage isolation, report battery pack status to the FCS, and / or detect and respond to faults in the circuitry.

[0086] Figure 2C and Figure 2DExemplary arrangements of circuitry and switches are detailed to provide charging flexibility by allowing auxiliary loads and / or electric motors and actuators to be energized or de-energized during charging. For example, as... Figure 2C As shown, in a switch-off / closed arrangement, the battery pack can be charged while the remaining HVPS circuitry remains disconnected. Charging switches K6 (positive) and K7 (negative) can close to allow the battery pack to charge. Simultaneously, main switches K1 and K2, as well as pre-charge switches K3 and K5, can be opened to prevent energizing the remaining HVPS circuitry. Furthermore, as... Figure 2D As shown, in a switch-off / closed arrangement, the battery pack can be charged while the auxiliary load is connected but the electric motor and actuator remain disconnected. Charging switch devices K6 (positive) and K7 (negative) can be closed to allow battery pack 120 to charge. Furthermore, main switch devices K1 and K2 can remain closed (after pre-charging) and K4 can remain open. In some embodiments, the BMU can coordinate with the charging control unit (described in detail below) to open and close the charging switch devices (e.g., K6 (positive) and / or K7 (negative)) based on the detected battery pack charge level and the desired charge level.

[0087] exist Figure 2E In the "Flight" mode, high-voltage power is supplied to the electric motor and / or tilt actuator. The BMU can command switch K5 to close, allowing pre-charging of the electric motor and / or tilt actuator. In some embodiments, once the BMU determines that pre-charging is complete (e.g., a set time has elapsed or the voltage or stored charge has risen to a threshold level), the BMU can open K5 and close K4, allowing high-voltage power to be fed to the electric motor and / or tilt actuator. Furthermore, the BMU can manage the main bus (e.g., detect measurements and react accordingly), estimate HV isolation, and detect faults. In some embodiments, certain fault responses are disabled in Flight mode. For example, even if a fault is detected (e.g., the isolation monitor detects a ground current), the switch will remain closed and power will remain connected.

[0088] Figure 3AA diagram of an aircraft circuitry consistent with embodiments of this disclosure is shown. A person (e.g., a pilot, technician) can select a desired mode via a mode switch 316. In some embodiments, the mode switch 316 is located on a pilot's instrument panel and / or control panel. In some embodiments, the mode switch 316 may be a device remote from the aircraft. For example, the mode switch 316 may include a transmitter to transmit the mode selection to the aircraft via a wireless link (e.g., a radio link). In some embodiments, the mode switch 316 may be a physical switch, knob, button, joystick, and / or any structure configured to be moved by a user. In some embodiments, the mode switch 316 requires force to unlock the joystick before changing the mode to avoid accidental mode changes. In some embodiments, the mode switch 316 may be a user interface element provided on a pilot's display screen or control panel. In some embodiments, the mode switch 316 may be a processor capable of receiving manual selection and / or voice commands from the pilot requesting a mode change. The mode switch 316 may include any mechanism that allows the pilot to select a desired operating mode. In some embodiments, mode switch 316 and / or associated panels for mode switch 316 (e.g., one or more controllers, one or more processors, etc.) may include physical constraints or logic that provide restrictions on mode transitions. For example, in some embodiments, mode switch 316 may include one or more of the following restrictions: allowing only transitions from off mode 201 to maintenance mode 202, maintenance mode 202 to ground mode 203, ground mode 203 to flight mode 204, flight mode 204 to ground mode 203, ground mode 203 to maintenance mode 202, and maintenance mode 202 to off mode 201.

[0089] In some embodiments, the selected mode is transmitted to a switching device in the LV power distribution box 314 (LV PDB). In some embodiments, the switching device is separate from the LV PDB 314. The switching device may be a relay (e.g., a single-pole single-throw relay), a transistor, a contactor, a controller, or any other device capable of switching power on and off. The switching device may allow power to flow from the battery pack to the CCU 312 in maintenance mode 202 and ground mode 203. In some embodiments, the selected mode is transmitted to a system control unit (SCU) 322 for controlling low-voltage power distribution. For example, in some embodiments, the selected mode may be transmitted to the SCU 322 when the ground unit provides low-voltage power to control low-voltage power distribution in flight mode 204, ground mode 203, and / or maintenance mode 202. In some embodiments, the selected mode may also be transmitted to the FCS 318. In some embodiments, mode selection is transmitted directly from the mode switch 316 to the FCS 318. For example, in some embodiments, when the selected mode is flight mode 204, mode switch 316 sends the selected mode to FCS 318 and / or one or more flight control computers. In some embodiments, mode and / or status selection is transmitted to FCS 318 via CCU 312 and / or other components. For example, in some embodiments, CCU 312 and / or other components may transmit the mode and / or status when the aircraft is in ground mode 203, when low pressure is provided by a ground unit, and / or when the aircraft is in maintenance mode. In other embodiments, mode switch 216 sends signals directly to FCS 318 in all aircraft modes and / or states.

[0090] As referenced above Figures 2B-2E In detail, the high-voltage power distribution 300 may include a combination of switching devices and / or fuses. For example, the high-voltage power distribution 300 may include a main contactor, an electric motor contactor, a charging contactor, a pre-charge relay, pyro-fuses, and a thermal fuse. Furthermore, as shown, the high-voltage power distribution 300 can supply power to various HV loads 308 (e.g., electric motors, low-voltage systems, equipment) and can supply power to one or more paired battery packs 306 (e.g., via the methods described above). Figure 1J-Figure 1L The crossover line 130 shown in the diagram.

[0091] The Battery Management Unit (BMU) 302 can monitor the status of one or more battery packs and can communicate with various systems inside and outside the battery packs. For example, the BMU 302 can receive voltage, current, resistance, and temperature sensing signals from the battery stack assembly (and CMU 304) and / or HV distribution 300. In some embodiments, the BMU 302 performs calculations of the battery pack's State of Charge (SOC), State of Health (SOH), fault conditions (e.g., short circuit or overcurrent), State of Power (SOP), State of Energy (SOE), and State of Temperature (SOT). The BMU 302 also controls and monitors bus precharge, provides fuse and contactor commands, and communicates with various systems inside and outside the battery packs. In some embodiments, the BMU 302 can communicate with the flight control system 318. In some embodiments, the BMU 302 can receive power from the low-voltage system 320 included in subsystem 103. In some embodiments, the BMU 302 can receive power from one or more battery pack stacks (e.g., as referenced above). Figures 2B-2E (As shown).

[0092] In some embodiments, one or more high-voltage battery packs may power the charging control unit (CCU) 312. In some embodiments, only one high-voltage battery pack may power the CCU 312. In some embodiments, the CCU 312 may have a dedicated high-voltage circuitry and DC / DC converter separate from the high-voltage circuitry in subsystems 102 and 101. In some embodiments, power may be available to the LV PDB 314 regardless of the aircraft mode. In these embodiments, for example, the CCU 312 may be enabled in ground mode 203 and maintenance mode 202 even when subsystem 102, which includes a high-voltage converter for low-voltage power distribution, and subsystem 103, which includes a low-voltage power distribution system, are deactivated. As described above, in some embodiments, instead of the CCU 312, a separate controller (e.g., a central battery management unit and / or battery management system) may be powered and receive a selected mode via a switching device in the LV PDB 314. In some embodiments, the separate controller may comprise multiple controllers for redundancy. One or more controllers may include one or more power connections via a switching device. One or more controllers may perform communication with and / or control of different components of the aircraft. In some embodiments, one or more controllers perform any functions disclosed to be performed by the CCU and / or BMU.

[0093] Figure 3BAnother diagram of an aircraft circuitry consistent with embodiments of this disclosure is shown. As illustrated, subsystem 102 provides high-voltage power to one or more DC / DC converters (e.g., 328, 330, 332, 326) to feed power to subsystem 103, which includes a low-voltage power distribution circuitry and / or components. In some embodiments, one or more of the DC / DC converters may be alternative DC / DC converters (e.g., 326) that allow the aircraft to be powered and / or serve as a backup for other DC / DC converters. In some embodiments, in ground mode 203, one or more alternative DC / DC converters may feed power to subsystem 103, which includes a low-voltage power distribution circuitry and / or components.

[0094] In addition, the aircraft may include a low-voltage port assembly 334 that allows connection between subsystem 103 and a low-voltage ground power unit (LV GPU 324). In some embodiments, the low-voltage port assembly 334 may include one or more switching devices (e.g., relays, contactors, transistors, etc.) that allow low-voltage power to be connected to subsystem 103, which includes a low-voltage power distribution circuit system and / or components. In some embodiments, the low-voltage port assembly 334 may communicate with one or more components (e.g., SCU 332, CCU 312, etc.) when connected to low-voltage power from the LV GPU 324. In some embodiments, the LV GPU 324 may be connected in maintenance mode 202 to allow for aircraft maintenance (e.g., testing, software loading, etc.).

[0095] Figure 3C A diagram showing a mode switch 316 consistent with embodiments of this disclosure is provided. As described above, the selected mode is transmitted to switching devices (e.g., relays and / or transistors) in the power distribution box 314 and CCU 312. In ground mode 203 and maintenance mode 202, the switching devices in the power distribution box 314 are closed. Once the switching devices are closed, power can flow from the battery pack and DC / DC converter to power the CCU 312. Once powered, the CCU 312 receives the selected mode, ground mode 203, or maintenance mode 202, and communicates with the battery management unit 302 according to the selected mode. As described above, in some embodiments, instead of the CCU 312, a separate controller (e.g., a central battery management unit and / or battery management system) can be powered and receive the selected mode via the switching devices in the LV PDB 314.

[0096] Figure 3D Another figure shows a mode switch 316 consistent with an embodiment of this disclosure. (Refer to the above text.) Figure 3AAs described, different subsystems can be enabled and / or disabled in different modes. Figure 3A The diagram schematically illustrates several different components that can communicate and / or control in certain modes. In some embodiments, one or more subsystems of these subsystems can be directly enabled and / or disabled by mode switch 316. In some embodiments, one or more subsystems of these subsystems can be enabled and / or disabled via mode switch 316, which communicates with one or more other components. When off mode 201 is selected on mode switch 316, mode switch 316 can communicate with and / or control LV SCU 322 (e.g., directly or via other components) to disable the subsystem, as referenced above. Figure 2A As described above. When maintenance mode 202 is selected on mode switch 316, mode switch 316 can communicate with and / or control LV SCU 322 and / or CCU 312 (e.g., directly or via other components) to enable and / or disable subsystems, as referenced above. Figure 2A As described above. When ground mode 203 is selected on mode switch 316, mode switch 316 can communicate with and / or control (e.g., directly or via other components) LV SCU 322, CCU 312 and / or one or more power distribution boxes to enable and / or disable subsystems, as described above. Figure 2A As described above. When flight mode 204 is selected on mode switch 316, mode switch 316 can communicate with and / or control (e.g., directly or via other components) LV SCU 322, FCS 318 and / or one or more power distribution boxes to enable and / or disable subsystems, as described above. Figure 2A As described above, in some embodiments, instead of CCU 312, a separate controller (e.g., a central battery management unit and / or a battery management system) may receive the selected mode.

[0097] Figure 4A flowchart for switching aircraft modes, consistent with embodiments of this disclosure, is shown. As detailed in the flowchart, in some mode transitions, certain conditions must be met before the control subsystem transitions to the new selected mode. However, in some mode transitions, preconditions need not be met before the transition. In some embodiments, FCS 318 stores the calculated mode to which the aircraft transitions, while in other embodiments, FCS 318 does not store the calculated mode. At step 500, CCU 312 and / or FCS 318 may detect the selection of an aircraft mode on mode switch 316 (e.g., by a user at an interface). Aircraft modes may include: shutdown mode 201, maintenance mode 202, ground mode 203, and flight mode 204. At step 502, if applicable, the device may check whether conditions for transitioning the aircraft to the selected mode are met. At step 504, if the conditions are not met, the transition is prohibited and the subsystem remains in its current state. At step 506, if the condition is met (or if the condition is not required to change), the subsystem will be controlled according to the selected mode. For example, subsystems 101, 102, 103, 104, and / or 105 can be enabled or disabled based on the mode, as referenced above. Figure 2A As shown. In some embodiments, Figure 5 The outlined steps can be performed by CCU 312 and / or FCS 318. In some embodiments, the component performing the steps can vary based on which modes the aircraft is transitioning between. In some embodiments, FCS 318 can update the calculated modes. The following is a description of an example mode transition consistent with embodiments of this disclosure.

[0098] From shutdown mode 201 to maintenance mode 202: At step 500, the device (e.g., CCU 312) may detect a selection of a change from off mode 201 to maintenance mode 202 on mode switch 316 (e.g., by the user at the interface). In some embodiments, there are no preconditions for such a transition. At step 506, subsystem 105 is enabled. CCU 312 is powered on and has received a signal indicating that maintenance mode 202 has been selected.

[0099] From maintenance mode 202 to ground mode 203: At step 500, the device (e.g., CCU 312) may detect a selection of a change from maintenance mode 202 to ground mode 203 on mode switch 316 (e.g., by a user at an interface). CCU 312 may be powered (and / or remain powered) and receive a signal indicating that ground mode 203 has been selected. In some embodiments, CCU 312 may transmit the selected mode to one or more battery management units 302 and / or to flight control system (FCS) 318. At step 502, one or more BMUs 302 may determine whether the relevant loads satisfy the HVIL connectivity state. For example, one or more BMUs 302 may determine that the HVIL connectivity state is satisfied when a DC / DC converter is connected on both sides and / or other auxiliary loads (e.g., heaters, compressors, and / or other thermal regulation components) are properly connected. Furthermore, BMU 302 and / or CCU 312 can control the pre-charging of one or more battery packs and can confirm pre-charging completion before transitioning to ground mode 203 (e.g., a set time has elapsed or the voltage or stored charge has risen to a threshold level). If one or more conditions are not met, BMU 302 may not allow transition to ground mode 203. At step 504, if the conditions are not met, the aircraft may not transition to ground mode 203. At step 506, if the connectivity conditions are met, the aircraft may transition to ground mode 203. Subsystem 102, including high voltage supplied to a DC / DC converter for low-voltage power distribution, may be enabled. Subsystem 103, including a low-voltage power distribution system, may also be enabled. In some embodiments, in ground mode 203, low-voltage power will come from an alternative DC / DC converter.

[0100] In some embodiments, switching to ground mode 203 may include powering SCU 322, which then receives a signal from mode switch 216 and / or another component (e.g., CCU 312) indicating the selection of ground mode 203. In some embodiments, switching to ground mode 203 may include system control unit (SCU) 322 commanding one or more low-voltage power distribution boxes (LV PDBs) to enable power delivery to one or more low-voltage systems. For example, in some embodiments, SCU 322 may command three power distribution boxes connected to three low-voltage buses to allow the distribution of low-voltage power. In some embodiments, FCS 318 may change its calculated mode to ground mode. In some embodiments, after switching to ground mode 203 and / or changing the calculated mode to ground mode 203, FCS 318 will take over control of the aircraft (e.g., from CCU 312). FCS 318 will issue specific commands to each subsystem based on user input, sensor input, flight status, etc.

[0101] From Ground Mode 203 to Flight Mode 204: At step 500, a device (e.g., FCS 318) may detect (e.g., by a user at an interface) a selection to change from ground mode 203 to flight mode 204. In some embodiments, the selection of flight mode may be transmitted to a system control unit (SCU) 322, one or more LV power distribution boxes (LV PDBs), and flight control system 318. The CCU 312 will not be powered. In some embodiments, the CCU 312 will be de-energized by a switching device in the LV PDB 314 disconnecting the CCU 312 from power. At step 502, one or more conditions may be checked. In some embodiments, the FCS 318 and / or one or more other devices may check the battery pack temperature to confirm it is within the permissible range for flight. In some embodiments, the permissible range may be a temperature range that allows the aircraft to safely perform its next mission. In some embodiments, the FCS 318 and / or one or more other devices may confirm that no charger is connected to the aircraft. In some embodiments, the FCS 318 and / or one or more other devices may confirm that no coolant lines are connected to the aircraft. In some embodiments, FCS 318 and / or one or more other devices may confirm that no other plugs for powering the low-voltage power distribution system are connected to the aircraft. In some embodiments, BMU 302 and / or one or more other devices may check the HVIL connectivity status of the electric engine and tilt actuator. In some embodiments, BMU 302 may provide connectivity status information to FCS 318. At step 504, if any one of the conditions is not met, transition to flight mode 204 is prohibited. In some embodiments, the calculated mode of FCS 318 remains ground mode 203. At step 506, if all or one of the conditions are met, transition to flight mode 204 is permitted. Subsystem 101 is activated to provide high-voltage power to the electric engine and tilt actuator. In some embodiments, in flight mode 204, subsystem 103, which includes low-voltage power distribution, is fed from the main DC / DC converter, and an alternative DC / DC converter is used as a backup. In some embodiments, FCS 318 stores the calculated mode as flight mode 204. In some embodiments, the FCS 318 can communicate with one or more aircraft subsystems to indicate that the mode is flight mode 204.

[0102] From Flight Mode 204 to Ground Mode 203: At step 500, a device (e.g., CCU 312) may detect a selection of a change from flight mode 204 to ground mode 203 (e.g., by a user at an interface). As described above, a signal may be sent to LV PDB 314, and CCU 312 may be powered. In some embodiments, the mode selection may be transmitted to System Control Unit (SCU) 322 and / or Flight Control System 318. At step 502, FCS 318 and / or one or more other devices may confirm that the aircraft is on the ground. In some embodiments, this may involve detecting landing gear deployment and / or the aircraft being on the ground from one or more sensors associated with the landing gear (e.g., shock absorber sensors monitoring deflection). Alternatively or additionally, FCS 318 and / or one or more other devices may confirm that the aircraft is stationary (e.g., via GPS sensors, laser speed sensors, wheel speed sensors, wheel tooth sensors, etc.). Alternatively or additionally, the aircraft may be confirmed as stationary when the speed is below a set threshold or when the speed is zero. Alternatively, FCS 318 and / or one or more other devices may determine whether the aircraft is on the ground by determining whether the aircraft's altitude is below a threshold (e.g., using an altimeter). At step 504, if any one of the conditions is not met, transition to ground mode 203 is prohibited. At step 506, if all one or more conditions are met, transition to ground mode 203 is permitted. Subsystem 101 may be deactivated, and power to the electric motor and tilt actuators will be removed. In some embodiments, FCS 318 stores the calculated mode as ground mode 203.

[0103] From ground mode 203 to maintenance mode 202: At step 500, the device (e.g., CCU 312) may detect a selection of a change from ground mode 203 to maintenance mode 202 (e.g., by the user at the interface). As described above, a signal may be sent to LV PDB 314, and charging control unit 312 may be powered. In some embodiments, there are no preconditions for such a switch. At step 506, subsystem 103 is deactivated, and the low-voltage power distribution fed from the battery pack is removed. Furthermore, subsystem 102 may be deactivated, and the high-voltage power supplied to the DC / DC converter is removed. In some embodiments, FCS 318 stores the calculated mode as maintenance mode 202.

[0104] From maintenance mode 202 to shutdown mode 201: At step 500, the device (e.g., SCU 322) may detect a selection of a change from maintenance mode 202 to shutdown mode 201 (e.g., at the user interface). In some embodiments, there are no preconditions for implementing this transition. At step 406, subsystem 105, including power supplied to CCU 312, is deactivated.

[0105] Figure 5 Exemplary transitions between states within a pattern consistent with embodiments of this disclosure are illustrated. As detailed in the flowcharts, in some state transitions, CCU 312 and / or FCS 318 may not control the subsystem in the new state unless certain conditions are met. However, in some state transitions, preconditions need not be met before the transition. In some embodiments, the steps of FIG. 6 may be performed by CCU 312 and / or FCS 318. In some embodiments, the components performing the steps may vary based on which states the aircraft is transitioning between. At step 600, the aircraft may detect events such as user selection and / or connected power. In some embodiments, the aircraft's state may include a ground unit charging state, a low-voltage supply state from the ground unit, and / or a maintenance state. At step 602, if applicable, the device may check whether conditions for transitioning the aircraft to the new state are met. At step 604, if the conditions are not met, the transition is prohibited and the subsystem remains in its current state. At step 606, if the conditions are met (or if the transition does not require conditions), the subsystem is controlled according to the new state. For example, subsystems 101, 102, 103, 104, and / or 105 can be enabled or disabled based on the selected state. In some embodiments, the FCS 318 can store new states.

[0106] Switching to charging mode: At step 600, in maintenance mode 202 and ground mode 203, CCU 312 can detect that a charger (e.g., a power source) is connected to the associated charging port on the aircraft. At step 602, CCU 312 can communicate with the battery pack (e.g., via BMU 302) to detect the voltage and / or current in the battery pack and confirm whether they are within predetermined limits to allow charging. At step 604, if the battery pack voltage and / or current conditions are not met, CCU 312 can transmit details of the problem (e.g., the type of problem, such as insufficient current or voltage, one or more associated battery packs, etc.) to the ground charging equipment. The ground charging equipment can transmit the problem via a display, computer, laptop computer, iPad, mobile device, or any other device capable of transmitting information to charging service personnel.

[0107] At step 606, if all or one of the conditions are met, CCU 312 and / or charging subsystem 310 may engage a latch to prevent the high-voltage power plug from disconnecting. CCU 312 may enable subsystem 104, which includes a charging bus and associated devices. For example, CCU 312 may control switching devices to allow power flow between the ground charging equipment and the battery pack. Furthermore, CCU 312 and / or charging subsystem 310 may send a request for power to the ground charging equipment. In some embodiments, flight control system (FCS) 318 may update its current state to charging.

[0108] Low-pressure status from the ground unit: At step 600, in maintenance mode 202 and ground mode 203, the low-voltage CPU and / or one or more other devices can detect that the ground unit is connected. In some embodiments, the ground unit can supply high-voltage power reduced from a DC / DC converter (e.g., an alternative DC / DC converter). In other embodiments, the ground unit can directly supply low-voltage power. At step 602, the low-voltage CPU and / or one or more other devices can verify whether the power quality meets the aircraft's requirements. At step 604, if the power quality condition is not met, the aircraft will not switch to allowing low-voltage power from the ground unit. At step 606, if the power quality condition is met, the aircraft will switch to allowing low-voltage power from the ground unit. In some embodiments, the system control unit (SCU) 322 can command one or more low-voltage power distribution boxes to enable power to one or more low-voltage components. For example, in some embodiments, the SCU 322 can command three power distribution boxes connected to three low-voltage buses to allow low-voltage power to be fed to various low-voltage components. In some embodiments, the flight control system (FCS) 318 may update its current state to indicate low-voltage power supplied by the ground unit.

[0109] Maintenance status in maintenance mode 202: At step 600, in maintenance mode 202, CCU 312 and / or FCS 318 can detect the selection of a maintenance state to perform a test or software upload on the aircraft. In some embodiments, this selection can be made using mode switch 316, while in other embodiments, it can be made using a separate user input device. In some embodiments, the user input device may include a joystick that requires force to unlock before changing to a maintenance state to prevent accidental selection. In some embodiments, the maintenance switch may be a user interface element provided to the pilot on a display screen or control panel. In some embodiments, the maintenance switch may be a processor that receives manual selection and / or voice commands from the pilot requesting a maintenance state.

[0110] At step 602, if a maintenance state is selected (e.g., as detected by CCU 312), CCU 312 can transmit the selection to flight control system 318. In some embodiments, this transmission can be made via one or more battery management units 302. In some embodiments, CCU 312 can communicate directly with FCS 318. In some embodiments, there are no preconditions for transitioning to the maintenance mode maintenance state. At step 606, FCS 318 can allow maintenance of the aircraft in maintenance mode 202. For example, FCS 318 can control various subsystems to allow software loading, including the loading of software for battery management unit 302 (e.g., operations that FCS 318 might not allow in other modes). Furthermore, FCS 318 can control various subsystems to allow testing of low-voltage equipment. For example, FCS 318 can control testing of one or more flight control computers, aircraft flight control surfaces, motor controllers, pitch actuators, tilt actuators, battery management systems for battery packs, environmental control systems, sensors, and / or any other low-voltage systems on the aircraft. FCS 318 can store the maintenance mode state as the current state.

[0111] Maintenance status in ground mode 203: At step 600, in ground mode 203, CCU 312 and / or FCS 318 can detect a selection of a maintenance state (e.g., by a user at an interface) to perform a test or software upload on the aircraft. In some embodiments, this selection can be made using mode switch 316. As described above, in some embodiments, this selection is made by a separate device. For example, by a joystick that requires force to unlock before changing to a maintenance state to prevent accidental selection. In some embodiments, the maintenance switch can be a user interface element provided to the pilot on a display screen or control panel. In some embodiments, the maintenance switch can be a processor that receives manual selection and / or voice commands from the pilot requesting a maintenance state.

[0112] At step 602, CCU 312 may transmit the selection to flight control system 318. In some embodiments, this transmission may be made via one or more BMUs 302. In some embodiments, CCU 312 may communicate directly with FCS 318. Before allowing a transition to ground mode maintenance and / or before allowing testing of high-voltage equipment, BMU 302 may determine whether the HVIL connectivity state is met, thereby indicating that the battery pack is connected at both ends (e.g., both sides of a DC / DC converter). At step 604, if one or more conditions are not met, ground mode maintenance and / or testing may be disabled. At step 606, if all or one or more conditions are met, FCS 318 may allow ground mode maintenance. For example, FCS 318 may control various subsystems to allow software loading (e.g., operations that FCS 318 may not allow in other modes). Furthermore, FCS 318 may control various subsystems to allow testing of low-voltage equipment (e.g., operations that FCS 318 may not allow in other modes). For example, the FCS 318 can control one or more flight control computers, aircraft flight control surfaces, motor controllers, pitch actuators, tilt actuators, battery management systems for battery packs, environmental control systems, sensors, and / or any other systems on the aircraft that require testing of low-voltage electricity.

[0113] Furthermore, in ground mode 203, FCS 318 can allow testing of high-voltage equipment. For example, FCS 318 can control subsystems to allow the following operations: testing electric engines, associated electric propulsion units such as rotors, propellers, and / or steers; and testing tilting propeller systems, including linear actuators and / or rotary actuators. In some embodiments, FCS 318 can store maintenance status as the current state.

[0114] As discussed above, in some embodiments, a connectivity check may be performed before switching aircraft modes or states. A connectivity check may be performed to ensure that the relevant circuitry is intact and free of open circuits (e.g., breaks or gaps in the circuitry). In some embodiments, a connectivity check may involve placing a small voltage in the relevant circuitry and monitoring whether current flows through the circuit. If current is detected or the current meets an expected value, connectivity is confirmed. If no current is detected or the current does not meet an expected value, connectivity is not confirmed. As described above, if connectivity is not confirmed, the aircraft may be prevented from transitioning to the next mode or state. In some embodiments, if connectivity is not confirmed, an alarm may be provided indicating that a transition to the next mode or state may not be performed. In some embodiments, the alarm may indicate components and / or circuits for which connectivity cannot be confirmed. In some embodiments, the alarm may include displaying an image and / or text on a screen, turning on lights, or activating a sound.

[0115] In some embodiments, instead of and / or in addition to applying a small voltage to the relevant circuit to check its connectivity, the voltage supplied to the component can be compared with the voltage measured at the component. If a voltage is detected at the component and / or the voltage meets the expected value, the connectivity of the relevant circuit is confirmed. If no voltage is detected at the component and / or the voltage does not meet the expected value, the connectivity of the relevant circuit is not confirmed. As described above, if circuit connectivity is not confirmed, the aircraft may be prevented from transitioning to the next mode or state, and an alarm may be provided. In some embodiments, a connectivity check of a DC / DC converter involves applying a small voltage to the DC / DC converter to check its connectivity. In some embodiments, a connectivity check of an electric engine and / or tilt actuator involves determining whether a received voltage is detected or meets the expected value when the power supply provides voltage. The connectivity check can be performed by the battery management system of the battery pack. The results of the connectivity check can be transmitted to the charging control unit and / or the flight control system, depending on which component is performing mode or state switching verification.

[0116] The foregoing description has been presented for illustrative purposes. This description is not exhaustive and does not limit the invention to the precise forms or embodiments disclosed. Modifications and adaptations to the invention will be apparent to those skilled in the art upon consideration of the specification and the disclosed embodiments thereof.

[0117] The features and advantages of this disclosure are apparent from the detailed description, and therefore the appended claims are intended to cover all systems and methods falling within 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 multiple unless explicitly stated in the given context. Unless otherwise explicitly indicated, words such as “and” or “or” mean “and / or”. Furthermore, since various modifications and alterations will be readily made by studying this disclosure, it is not intended to limit this disclosure to the exact constructions and operations shown and described, and therefore all suitable modifications and equivalents falling within the scope of this disclosure may be employed.

[0118] Other embodiments will become apparent to those skilled in the art upon consideration of the description and the embodiments disclosed herein. The structures and circuit arrangements shown in the figures are intended for illustrative purposes only and are not intended to limit the specific arrangements and circuit configurations described and illustrated in the figures. The description and examples are also intended to be considered merely exemplary, and the true scope and spirit of the invention are indicated by the appended claims. The foregoing description has been presented for illustrative purposes. This description is not exhaustive and does not limit the invention to the precise forms or embodiments disclosed. Modifications and adaptations to the invention will become apparent to those skilled in the art upon consideration of the disclosed embodiments of the invention as described herein.

[0119] Additional aspects of this disclosure may be further described by the following terms: 1. A computer-implemented method for controlling the power distribution of an aircraft, the computer-implemented method comprising: The aircraft receives a selection of at least three aircraft operating modes from a user input device at the control circuitry within the aircraft; and Based on the selected operating mode, the power distribution within the aircraft is controlled via the control circuitry, wherein controlling the power distribution based on the selected operating mode includes: Based on the selected operating mode, the high-voltage power supplied to at least one electric propulsion unit and the high-voltage power supplied to at least one non-propulsion load are individually controlled via the control circuit.

[0120] 2. The computer-implemented method according to Clause 1, wherein the user input device is at least one of the following: a switch, a knob, a button, a joystick, a display screen element, or a voice receiver.

[0121] 3. A computer-implemented method according to Clause 1 or 2, wherein controlling the power distribution within the aircraft includes controlling at least one switching device.

[0122] 4. The computer-implemented method according to Clause 3, wherein the switching device is at least one of a contactor, a relay, or a transistor.

[0123] 5. A computer-implemented method according to any one of clauses 1 to 3, wherein in one of the at least three aircraft modes, controlling the power distribution within the aircraft includes not providing high-voltage power to the at least one electric propulsion unit and not providing high-voltage power to the at least one non-propulsion load.

[0124] 6. A computer-implemented method according to any one of clauses 1 to 5, wherein: Controlling the power distribution within the aircraft based on the selected operating mode further includes controlling the power sent to the controller. The controller is configured to control the high-voltage power; and In one of the at least three aircraft modes, controlling the power distribution within the aircraft includes supplying power to the controller.

[0125] 7. A computer-implemented method according to any one of clauses 1 to 6, wherein in one of the at least three aircraft modes, controlling the power distribution within the aircraft includes not providing high-voltage power to the at least one electric propulsion unit and providing high-voltage power to the at least one non-propulsion load.

[0126] 8. The computer-implemented method according to Clause 7, wherein the computer-implemented method further comprises: Upon receiving the selection of the chosen mode and before controlling the power distribution within the aircraft, it is determined whether one or more conditions are met. The conditions being met include at least one of the following: meeting the pre-charge requirement or meeting the connectivity requirement.

[0127] 9. A computer-implemented method according to any one of clauses 1 to 8, wherein in one of the at least three aircraft modes, controlling the power distribution within the aircraft includes providing high-voltage power to the at least one electric propulsion unit and to the at least one non-propulsion load.

[0128] 10. The computer-implemented method according to Clause 9, wherein the computer-implemented method further comprises: Upon receiving the selection of the chosen mode and before controlling the power distribution within the aircraft, it is determined whether one or more conditions are met. Meeting one or more of the conditions includes at least one of the following: meeting connectivity requirements for the at least one electric propulsion unit, meeting pre-charge requirements, the battery pack being within a safe temperature range, not having a coolant line connected, not having a charger connected, or not having a low-voltage plug connected.

[0129] 11. The computer-implemented method according to Clause 9, wherein the computer-implemented method further comprises: Upon receiving a selection for a new mode after selecting one of the chosen modes, and before controlling the power distribution within the aircraft based on the new selected mode, it is determined whether one or more conditions are met, wherein meeting the one or more conditions includes at least one of the following: the aircraft is stationary, the aircraft speed is below a threshold, or the aircraft landing gear is deployed; and When one or more of the conditions are determined to be met, the power distribution within the aircraft is controlled via the control circuit based on the new selected operating mode.

[0130] 12. The computer-implemented method according to any one of clauses 1 to 11, wherein the computer-implemented method further comprises: The system detects whether the charger is connected in one of the at least three aircraft modes, but not in another of the at least three aircraft modes. When the charger is detected to be connected, determine whether one or more conditions for charging are met; and Charging is enabled when one or more of the conditions are met.

[0131] 13. The computer-implemented method according to any one of clauses 1 to 12, wherein the computer-implemented method further comprises: Detect whether the low-voltage power supply is connected in one of the at least three aircraft modes, but not connected in another of the at least three aircraft modes; and When the low-voltage power supply is detected to be connected, determine whether one or more conditions for accepting low-voltage power are met; When it is determined that one or more of the conditions are met, control the low-voltage power supplied to at least one aircraft subsystem.

[0132] 14. The computer-implemented method according to any one of clauses 1 to 13, wherein the computer-implemented method further comprises: Upon receiving one of the at least three aircraft modes and before controlling the power distribution within the aircraft, it is determined whether one or more conditions are met, wherein meeting the one or more conditions includes detecting a low-voltage current through the non-propulsion load.

[0133] 15. A computer-implemented method according to any one of clauses 1 to 14, wherein the computer-implemented method further comprises: Upon receiving one of the at least three aircraft modes and before controlling the power distribution within the aircraft, it is determined whether one or more conditions are met, wherein meeting the one or more conditions includes verifying the voltage on the at least one electric propulsion unit.

[0134] 16. A computer-implemented method according to any one of clauses 1 to 15, wherein the user input device is located remotely from the aircraft.

[0135] 17. A computer-implemented method according to any one of clauses 1 to 16, wherein the control circuit is at least one of a charging control unit or a flight control system.

[0136] 18. A computer-implemented method according to any one of clauses 1 to 17, wherein the computer-implemented method further comprises: In one of the at least three operating modes, the control circuit is a non-flight control system, and the non-flight control system receives the selection of the aircraft mode; and In another of the at least three operating modes, the control circuit is a flight control system, and the flight control system receives the selection of the aircraft mode.

[0137] 19. A computer-implemented method according to any one of clauses 1 to 18, wherein controlling the power distribution within the aircraft includes distributing high-voltage power in different ways in each of the at least three aircraft operating modes.

[0138] 20. A computer-implemented method according to Clause 19, wherein controlling the power distribution within the aircraft includes distributing low-voltage power in the same manner in at least two of the at least three aircraft operating modes.

[0139] 21. A computer-implemented method according to Clause 19, wherein controlling the power distribution within the aircraft includes distributing low-voltage power in different ways in at least two of the at least three aircraft operating modes.

[0140] 22. A computer-implemented method according to any one of clauses 1 to 21, wherein controlling the high-voltage power supplied to the at least one electric propulsion unit includes controlling the high-voltage power supplied to at least two electric propulsion units.

[0141] 23. A computer-implemented method according to any one of clauses 1 to 22, wherein controlling the high-voltage power supplied to the at least one electric propulsion unit includes controlling the high-voltage power supplied to all electric propulsion units located at the leading edge of the wing of the aircraft or the trailing edge of the wing of the aircraft.

[0142] 24. A computer-implemented method according to any one of clauses 1 to 23, wherein controlling the high-voltage power supplied to the at least one electric propulsion unit includes controlling the high-voltage power supplied to all electric propulsion units on the aircraft.

[0143] 25. An electrical control system for an aircraft, the electrical control system comprising at least one processor configured to execute instructions to cause the system to perform a method according to any one of clauses 1 to 24.

[0144] 26. An aircraft comprising an electrical control system pursuant to Clause 25.

[0145] 27. A computer-readable medium storing instructions that, when executed by at least one processor, cause the at least one processor to perform a method according to any one of clauses 1 to 24.

[0146] 28. A system for an aircraft, the system comprising: Battery Management Unit; One or more battery cells, the one or more battery cells being configured to supply high-voltage power; At least one first switching device, the at least one first switching device being configured to enable and disable the power supply from the charging port to the one or more battery cells; At least one second switching device, the at least one second switching device being configured to enable and disable the power supply from the one or more battery cells to the non-propulsion load; and At least one third switching device, the at least one third switching device being configured to enable and disable the power supply from the one or more battery cells to the electric propulsion unit of the aircraft; The battery management unit controls the first switching device, the second switching device, and the third switching device based on the selection of one of at least three aircraft operating modes received from the user input device.

[0147] 29. The system pursuant to Clause 28, wherein the battery management unit is configured to command the at least one first switching device, the at least one second switching device, and the at least one third switching device to deactivate the power supply in one of the at least three aircraft modes.

[0148] 30. A system according to Clause 28 or 29, wherein in one of the at least three aircraft modes, the battery management unit is configured to command the at least one first switching device to enable power supply from the charging port when it is determined that the charger is connected to the charging port, while the second switching device and the third switching device continue to disable power supply.

[0149] 31. A system according to any one of clauses 28 to 30, wherein the non-propulsion load includes a converter configured to reduce power to feed power to at least one low-voltage system of the aircraft.

[0150] 32. The system according to Clause 31, wherein in one of the at least three aircraft modes, the battery management unit is configured to command the at least one second switching device to enable power supply to the converter and the at least one low-voltage system.

[0151] 33. The system according to Clause 32, wherein in one of the at least three aircraft modes, the battery management unit is configured to command the at least one first switching device to enable power supply from the charging port when it is determined that the charger is connected, while the second switching device enables power supply and the third switching device disables power supply.

[0152] 34. The system according to any one of clauses 28 to 33, wherein the at least one second switching device is further configured to enable power supplied to the cross-connect line to provide backup power to the second battery cell.

[0153] 35. The system according to Clause 32, said system further includes: A first pre-charge resistor, configured to pre-charge the converter; and A first precharge switch, connected in series with a first precharge resistor, wherein in one of the at least three aircraft modes, the battery management unit is configured to close the first precharge switch and precharge the converter before commanding the at least one second switching device to enable power supply to the converter and the at least one low-voltage system.

[0154] 36. A system according to any one of clauses 28 to 35, wherein the battery management unit is configured to command the at least one third switching device to enable power supply from the one or more battery cells to the electric propulsion unit of the aircraft when the mode from the user input device is one of the at least three aircraft modes.

[0155] 37. The system pursuant to any one of clauses 28 to 36, A second pre-charge resistor, configured to pre-charge the electric propulsion unit; and A second pre-charge switch, connected in series with a second pre-charge resistor, wherein when operating in one of the selected modes, the battery management unit is configured to close the second pre-charge switch and pre-charge the electric propulsion unit before commanding the at least one third switching device to enable power supply to the electric propulsion unit of the aircraft.

[0156] 38. A system for an aircraft, the system comprising: User input device, the user input device being configured to receive input indicating an operating mode; A power switching device configured to provide power to the controller upon receiving a signal from the user input device, and A controller, configured to control power supplied to one or more subsystems of the aircraft; and The user input device is configured as follows: When the received input indicates the first operating mode, the signal is not sent to the power switching device; When the received input indicates a second operating mode, the signal is sent to the power switching device; When the received input indicates a third operating mode, the signal is sent to the power switching device; and When the received input indicates the fourth operating mode, the signal is not sent to the power switching device.

[0157] 39. A system pursuant to Clause 38, wherein the power switching device comprises at least one of the following: a relay, a transistor, a contactor, or a controller.

[0158] 40. In a system pursuant to Clause 39, the relays mentioned therein include single-pole single-throw relays.

[0159] 41. The system according to any one of clauses 38 to 40, the system further comprising a battery pack configured to provide high-voltage power to one or more electric propulsion units of the aircraft, wherein: The battery pack provides power to the power switching device; and The converter reduces the high-voltage power of the battery pack to power the power switching device.

[0160] 42. The system according to any one of clauses 38 to 41, wherein the system further includes the flight control system of the aircraft, wherein: When the received input indicates the first operating mode, the user input device does not provide a signal to the flight control system; and When the received input indicates the fourth operating mode, the user input device provides a signal to the flight control system.

[0161] 43. A system according to Clause 42, wherein when the received input indicates the second operating mode and the third operating mode, the user input device does not provide a signal to the flight control system.

[0162] 44. The system pursuant to Clause 42, wherein: When the received input indicates the second operating mode and the third operating mode, the controller provides the aircraft battery pack with information regarding the operating configuration; and When the received input indicates the fourth operating mode, the flight control system provides the aircraft battery pack with information about the operating configuration.

[0163] 45. The system according to Clause 44, wherein the flight control system is configured to provide information to the aircraft battery pack to supply power to one or more electric propulsion units when the received input indicates the fourth operating mode.

[0164] 46. ​​The system according to Clause 45, wherein the controller is configured to provide information to the aircraft battery pack to supply power to one or more low-voltage systems when the received input indicates the second operating mode.

[0165] 47. A system pursuant to any one of clauses 38 to 46, wherein: The controller is configured to detect the charging plug; and The controller is configured to control the aircraft battery pack during charging when the charging plug is detected.

[0166] 48. The system pursuant to any one of clauses 38 to 47, When the received input indicates the first operating mode, the aircraft is powered off; When the received input indicates the second operating mode, the aircraft is controlled by the controller to receive at least one of the following: charging information or test information; When the received input indicates the third operating mode, the aircraft is controlled by the controller to power at least one low-voltage system; and When the received input indicates the fourth operating mode and the controller remains unpowered, the aircraft is controlled by the aircraft's flight control system to provide power to at least one electric propulsion unit.

Claims

1. A computer-implemented method for controlling the power distribution of an aircraft, the computer-implemented method comprising: The aircraft receives a selection of one of at least three aircraft operating modes from a user input device at the control circuitry within the aircraft. as well as Based on the selected operating mode, the power distribution within the aircraft is controlled via the control circuitry, wherein controlling the power distribution based on the selected operating mode includes: Based on the selected operating mode, the high-voltage power supplied to at least one electric propulsion unit and the high-voltage power supplied to at least one non-propulsion load are individually controlled via the control circuit.

2. The computer-implemented method according to claim 1, wherein the user input device is at least one of the following: a switch, a knob, a button, a joystick, a display screen element, or a voice receiver.

3. The computer-implemented method according to claim 1 or 2, wherein controlling the power distribution within the aircraft includes controlling at least one switching device.

4. The computer-implemented method according to claim 3, wherein the switching device is at least one of a contactor, a relay, or a transistor.

5. The computer-implemented method according to any one of claims 1 to 3, wherein in one of the at least three aircraft modes, controlling the power distribution within the aircraft includes not providing high-voltage power to the at least one electric propulsion unit and not providing high-voltage power to the at least one non-propulsion load.

6. The computer-implemented method according to any one of claims 1 to 5, wherein: Controlling the power distribution within the aircraft based on the selected operating mode further includes controlling the power sent to the controller. The controller is configured to control the high-voltage power; and In one of the at least three aircraft modes, controlling the power distribution within the aircraft includes supplying power to the controller.

7. The computer-implemented method according to any one of claims 1 to 6, wherein in one of the at least three aircraft modes, controlling the power distribution within the aircraft includes not providing high-voltage power to the at least one electric propulsion unit and providing high-voltage power to the at least one non-propulsion load.

8. The computer-implemented method according to claim 7, further comprising: Upon receiving the selection of the chosen mode and before controlling the power distribution within the aircraft, it is determined whether one or more conditions are met. The conditions being met include at least one of the following: meeting the pre-charge requirement or meeting the connectivity requirement.

9. The computer-implemented method according to any one of claims 1 to 8, wherein in one of the at least three aircraft modes, controlling the power distribution within the aircraft includes providing high-voltage power to the at least one electric propulsion unit and to the at least one non-propulsion load.

10. The computer-implemented method according to claim 9, further comprising: Upon receiving the selection of the chosen mode and before controlling the power distribution within the aircraft, it is determined whether one or more conditions are met. Meeting one or more of the conditions includes at least one of the following: meeting connectivity requirements for the at least one electric propulsion unit, meeting pre-charge requirements, the battery pack being within a safe temperature range, not having a coolant line connected, not having a charger connected, or not having a low-voltage plug connected.

11. The computer-implemented method of claim 9, further comprising: Upon receiving a selection for a new mode after selecting one of the chosen modes, and before controlling the power distribution within the aircraft based on the new selected mode, it is determined whether one or more conditions are met, wherein meeting the one or more conditions includes at least one of the following: the aircraft is stationary, the aircraft speed is below a threshold, or the aircraft landing gear is deployed; and When one or more of the conditions are determined to be met, the power distribution within the aircraft is controlled via the control circuit based on the new selected operating mode.

12. The computer-implemented method according to any one of claims 1 to 11, wherein the computer-implemented method further comprises: The system detects whether the charger is connected in one of the at least three aircraft modes, but not in another of the at least three aircraft modes. When the charger is detected to be connected, determine whether one or more conditions for charging are met; as well as Charging is enabled when one or more of the conditions are met.

13. A computer-implemented method according to any one of claims 1 to 12, the computer-implemented method further comprising: Detect whether the low-voltage power supply is connected in one of the at least three aircraft modes, but not connected in another of the at least three aircraft modes; and When the low-voltage power supply is detected to be connected, determine whether one or more conditions for accepting low-voltage power are met; When it is determined that one or more of the conditions are met, control the low-voltage power supplied to at least one aircraft subsystem.

14. The computer-implemented method according to any one of claims 1 to 13, wherein the computer-implemented method further comprises: Upon receiving one of the at least three aircraft modes and before controlling the power distribution within the aircraft, it is determined whether one or more conditions are met, wherein meeting the one or more conditions includes detecting a low-voltage current through the non-propulsion load.

15. The computer-implemented method according to any one of claims 1 to 14, wherein the computer-implemented method further comprises: Upon receiving one of the at least three aircraft modes and before controlling the power distribution within the aircraft, it is determined whether one or more conditions are met, wherein meeting the one or more conditions includes verifying the voltage on the at least one electric propulsion unit.

16. The computer-implemented method according to any one of claims 1 to 15, wherein the user input device is located remotely from the aircraft.

17. The computer-implemented method according to any one of claims 1 to 16, wherein the control circuit is at least one of a charging control unit or a flight control system.

18. A computer-implemented method according to any one of claims 1 to 17, the computer-implemented method further comprising: In one of the at least three operating modes, the control circuit is a non-flight control system, and the non-flight control system receives the selection of the aircraft mode; and In another of the at least three operating modes, the control circuit is a flight control system, and the flight control system receives the selection of the aircraft mode.

19. The computer-implemented method according to any one of claims 1 to 18, wherein controlling the power distribution within the aircraft includes distributing high-voltage power in different ways in each of the at least three aircraft operating modes.

20. The computer-implemented method of claim 19, wherein controlling the power distribution within the aircraft comprises distributing low-voltage power in the same manner in at least two of the at least three aircraft operating modes.

21. The computer-implemented method of claim 19, wherein controlling the power distribution within the aircraft includes distributing low-voltage power in different ways in at least two of the at least three aircraft operating modes.

22. The computer-implemented method according to any one of claims 1 to 21, wherein controlling the high-voltage power supplied to the at least one electric propulsion unit includes controlling the high-voltage power supplied to at least two electric propulsion units.

23. The computer-implemented method according to any one of claims 1 to 22, wherein controlling the high-voltage power supplied to the at least one electric propulsion unit includes controlling the high-voltage power supplied to all electric propulsion units located on the leading edge or trailing edge of the wing of the aircraft.

24. The computer-implemented method according to any one of claims 1 to 23, wherein controlling the high-voltage power supplied to the at least one electric propulsion unit includes controlling the high-voltage power supplied to all electric propulsion units on the aircraft.

25. An electrical control system for an aircraft, the electrical control system comprising at least one processor configured to execute instructions to cause the system to perform the method according to any one of claims 1 to 24.

26. An aircraft comprising the electrical control system according to claim 25.

27. A computer-readable medium storing instructions that, when executed by at least one processor, cause the at least one processor to perform the method according to any one of claims 1 to 24.

28. A system for an aircraft, the system comprising: Battery Management Unit; One or more battery cells, the one or more battery cells being configured to supply high-voltage power; At least one first switching device, the at least one first switching device being configured to enable and disable the power supply from the charging port to the one or more battery cells; At least one second switching device, the at least one second switching device being configured to enable and disable the power supply from the one or more battery cells to the non-propulsion load; as well as At least one third switching device, the at least one third switching device being configured to enable and disable the power supply from the one or more battery cells to the electric propulsion unit of the aircraft; The battery management unit controls the first switching device, the second switching device, and the third switching device based on the selection of one of at least three aircraft operating modes received from the user input device.

29. The system of claim 28, wherein the battery management unit is configured to command the at least one first switching device, the at least one second switching device, and the at least one third switching device to deactivate power supply in one of the at least three aircraft modes.

30. The system of claim 28 or 29, wherein in one of the at least three aircraft modes, the battery management unit is configured to command the at least one first switching device to enable power supply from the charging port when it is determined that the charger is connected to the charging port, while the second switching device and the third switching device continue to disable power supply.

31. The system according to any one of claims 28 to 30, wherein the non-propulsion load comprises a converter configured to reduce power to feed power to at least one low-voltage system of the aircraft.

32. The system of claim 31, wherein in one of the at least three aircraft modes, the battery management unit is configured to command the at least one second switching device to enable power supply to the converter and the at least one low-voltage system.

33. The system of claim 32, wherein in one of the at least three aircraft modes, the battery management unit is configured to command the at least one first switching device to enable power supply from the charging port when it is determined that the charger is connected, while the second switching device enables power supply and the third switching device disables power supply.

34. The system according to any one of claims 28 to 33, wherein the at least one second switching device is further configured to enable power supplied to the cross-connect line to provide backup power to the second battery cell.

35. The system of claim 32, further comprising: A first pre-charge resistor is configured to pre-charge the converter; as well as A first precharge switch, connected in series with a first precharge resistor, wherein in one of the at least three aircraft modes, the battery management unit is configured to close the first precharge switch and precharge the converter before commanding the at least one second switching device to enable power supply to the converter and the at least one low-voltage system.

36. The system according to any one of claims 28 to 35, wherein the battery management unit is configured to command the at least one third switching device to enable power supply from the one or more battery cells to the electric propulsion unit of the aircraft when the mode from the user input device is one of the at least three aircraft modes.

37. The system according to any one of claims 28 to 36, the system further comprising: A second pre-charge resistor is configured to pre-charge the electric propulsion unit; as well as A second pre-charge switch, connected in series with a second pre-charge resistor, wherein when operating in one of the selected modes, the battery management unit is configured to close the second pre-charge switch and pre-charge the electric propulsion unit before commanding the at least one third switching device to enable power supply to the electric propulsion unit of the aircraft.

38. A system for an aircraft, the system comprising: User input device, the user input device being configured to receive input indicating an operating mode; A power switching device configured to provide power to the controller upon receiving a signal from the user input device, and A controller, configured to control power supplied to one or more subsystems of the aircraft; and The user input device is configured as follows: When the received input indicates the first operating mode, the signal is not sent to the power switching device; When the received input indicates a second operating mode, the signal is sent to the power switching device; When the received input indicates a third operating mode, the signal is sent to the power switching device; as well as When the received input indicates the fourth operating mode, the signal is not sent to the power switching device.

39. The system of claim 38, wherein the power switching device comprises at least one of the following: a relay, a transistor, a contactor, or a controller.

40. The system of claim 39, wherein the relay comprises a single-pole single-throw relay.

41. The system according to any one of claims 38 to 40, the system further comprising a battery pack configured to provide high-voltage power to one or more electric propulsion units of the aircraft, wherein: The battery pack provides power to the power switching device; and The converter reduces the high-voltage power of the battery pack to power the power switching device.

42. The system according to any one of claims 38 to 41, the system further comprising the flight control system of the aircraft, wherein: When the received input indicates the first operating mode, the user input device does not provide a signal to the flight control system; and When the received input indicates the fourth operating mode, the user input device provides a signal to the flight control system.

43. The system of claim 42, wherein when the received input indicates the second operating mode and the third operating mode, the user input device does not provide a signal to the flight control system.

44. The system according to claim 42, wherein: When the received input indicates the second operating mode and the third operating mode, the controller provides the aircraft battery pack with information regarding the operating configuration; and When the received input indicates the fourth operating mode, the flight control system provides the aircraft battery pack with information about the operating configuration.

45. The system of claim 44, wherein the flight control system is configured to provide information to the aircraft battery pack to supply power to one or more electric propulsion units when the received input indicates the fourth operating mode.

46. ​​The system of claim 45, wherein the controller is configured to provide information to the aircraft battery pack to supply power to one or more low-voltage systems when the received input indicates the second operating mode.

47. The system according to any one of claims 38 to 46, wherein: The controller is configured to detect the charging plug; and The controller is configured to control the aircraft battery pack during charging when the charging plug is detected.

48. The system according to any one of claims 38 to 47, When the received input indicates the first operating mode, the aircraft is powered off; When the received input indicates the second operating mode, the aircraft is controlled by the controller to receive at least one of the following: charging information or test information; When the received input indicates the third operating mode, the aircraft is controlled by the controller to power at least one low-voltage system; and When the received input indicates the fourth operating mode and the controller remains unpowered, the aircraft is controlled by the aircraft's flight control system to provide power to at least one electric propulsion unit.

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