High voltage battery architecture

Abstract

A power distribution system for an aircraft includes a plurality of electric propulsor units (EPUs) and includes a first paired battery pack unit including a first battery electrically connected to a second battery via a first high voltage bus. The first battery and the second battery are configured to provide power to a first set of EPUs and a second set of EPUs, respectively, of the plurality of EPUs. The system includes a second paired battery unit including a third battery electrically connected to a fourth battery via a second high voltage bus. The third battery and the fourth battery are configured to provide power to a third group of EPUs and a fourth group of EPUs, respectively, of the plurality of EPUs. The first high-voltage bus and the second high-voltage bus are electrically separated from each other.

Description

[0001] Cross - reference to related applications

[0002] This disclosure claims the benefit and priority of U.S. Provisional Application No. 63 / 383,660, filed on November 14, 2022, entitled "Systems and Methods for Improved Battery Assemblies for eVTOL Aircraft" (Attorney Docket No. 16163.6005 - 00000), the entire content of which is incorporated herein by reference for all purposes. Technical field

[0003] This disclosure generally relates to the field of electric aircraft. More specifically and without limitation, this disclosure relates to innovations in tilt - rotor aircraft using electric propulsion systems. Certain aspects of this disclosure generally relate to the configuration and control of high - voltage power systems for aircraft. Background art

[0004] Electric aircraft include battery packs that power various flight components, including electric propulsion units (EPUs) that enable flight. These battery packs are crucial for ensuring that the EPUs can provide lift and thrust support for the aircraft. Therefore, it is necessary to provide redundancy in the aircraft's power system to avoid single - point failures. It is also necessary to ensure that failure or fault conditions do not spread and damage other critical aircraft components. The disclosed high - voltage power system addresses these and other problems by connecting battery packs together in battery pack units, where each battery pack in the unit serves as a backup battery pack for the other battery packs. Further, each battery pack power supply is electrically isolated from other battery pack units.

[0005] Additionally, to ensure that the battery packs can power the EPU during flight, the battery packs need to be fully charged before takeoff. Therefore, it is necessary to charge the battery packs efficiently and effectively. The disclosed high - voltage power system addresses these and other problems by controlling the battery pack charge based on upcoming flight information, historical battery pack information, and the monitored status of the battery packs. The disclosed high - voltage power system also addresses this and other problems by providing single - point charging for multiple battery packs.

[0006] Finally, in a collision event, it is required that the first transponder can quickly and safely shut down the high-voltage power system. The disclosed high-voltage power system solves this problem by providing a low-voltage cut-off circuit connected to the battery pack. When it is detected that the first transponder has cut off the low-voltage cut-off circuit, the fuse of the battery pack is blown, and the high-voltage power system is no longer powered. The cut-off circuit can be arranged at the tail of the aircraft to provide isolation from the high-voltage line and increase the safety of the first transponder. Summary of the Invention

[0007] The present disclosure generally relates to a power system for an aircraft. One aspect of the present disclosure provides a power distribution system for an aircraft, the power distribution system including a plurality of electric propulsion units (EPUs). A system including a first paired battery pack unit includes a first battery electrically connected to a second battery via a first high-voltage bus, wherein the first battery is configured to supply power to a first group of the plurality of EPUs, and wherein the second battery is configured to supply power to a second group of the plurality of EPUs. The system includes a second paired battery pack unit, the second paired battery pack unit including a third battery electrically connected to a fourth battery via a second high-voltage bus, wherein the third battery is configured to supply power to a third group of the plurality of EPUs, and wherein the fourth battery is configured to supply power to a fourth group of the plurality of EPUs. Further, the first high-voltage bus and the second high-voltage bus are electrically separated from each other. Brief Description of the Drawings

[0008] Figure 1A An exemplary eVTOL aircraft consistent with an embodiment of the present disclosure is illustrated.

[0009] Figure 1B Another exemplary eVTOL aircraft consistent with an embodiment of the present disclosure is illustrated.

[0010] Figure 1C An electric engine 110 with two partial motors consistent with an embodiment of the present disclosure is illustrated.

[0011] Figure 1D A schematic diagram of a high-voltage power distribution system for an eVTOL aircraft consistent with an embodiment of the present disclosure is illustrated.

[0012] Figure 2A A circuit diagram of a high-voltage junction box (HVJB) consistent with an embodiment of the present disclosure is illustrated.

[0013] Figure 2B A schematic diagram of a high-voltage junction box (HVJB) consistent with an embodiment of the present disclosure is illustrated.

[0014] Figure 3 A schematic diagram of a charging port assembly (CPA) consistent with an embodiment of the present disclosure is illustrated.

[0015] Figure 4 Illustrated is a flowchart for detecting an emergency transponder consistent with an embodiment of the present disclosure.

[0016] Figure 5A Illustrated is a plan view of a cut-off circuit line arranged through the tail of an eVTOL aircraft consistent with an embodiment of the present disclosure.

[0017] Figure 5B Illustrated is a cross-sectional schematic view of a cut-off circuit line arranged through the tail of an eVTOL aircraft consistent with an embodiment of the present disclosure. Detailed Description

[0018] The present disclosure relates to components of an electric vertical takeoff and landing (eVTOL) aircraft mainly for unconventional aircraft. For example, the eVTOL aircraft of the present disclosure may be intended for frequent (e.g., more than 50 flights per working day), short-duration flights (e.g., each flight less than 100 miles) over, into, and out of populated areas. The aircraft may be intended to carry 4 - 6 passengers or commuters who expect a low-noise and low-vibration experience. Thus, it may be desirable for their components to be configured and designed to withstand frequent use without wear, require them to generate less heat and vibration, and require the aircraft to include mechanisms for effectively controlling and managing the heat or vibration generated by the components. Further, this may mean that several of these aircraft operate close to each other over crowded metropolitan areas. Accordingly, it may be desirable for their components to be configured and designed to generate low levels of noise inside and outside the aircraft and to have various safety and backup mechanisms. For example, for safety reasons, it may be desirable for the aircraft to be propelled by a distributed propulsion system, avoid the risk of single-point failures, and for the aircraft to be able to take off and land conventionally on a runway. Additionally, compared to traditional airport runways, it may be desirable for the aircraft to be able to take off and land vertically from relatively restricted spaces (e.g., vertical airports, parking lots, or driveways) while transporting approximately 4 - 6 passengers or commuters with luggage. These usage requirements may impose design constraints on the aircraft size, weight, operating efficiency (e.g., drag, energy use), which may affect the design and configuration of the aircraft components.

[0019] The disclosed embodiments provide new and improved configurations of aircraft components not observed in conventional aircraft, and / or design criteria for the identification of components different from conventional aircraft. Such alternative configurations and design criteria, in combination with addressing the disadvantages and challenges of conventional components, result in the embodiments disclosed herein for various configurations and designs of eVTOL aircraft components.

[0020] In some embodiments, the eVTOL aircraft of the present disclosure can be designed to be capable of vertical takeoff and landing as well as conventional takeoff and landing, wherein the distributed electric propulsion system enables vertical flight, forward flight, and transition flight. Thrust can be generated by supplying high-voltage electrical energy to the electric engines of the distributed electric propulsion system, and each electric engine can convert the high-voltage electrical energy into mechanical shaft power to rotate the propellers. Embodiments disclosed herein can relate to optimizing the energy density of the electric propulsion system. Embodiments can include electric engines connected to an on-board power source, which can include devices capable of storing energy (such as batteries or capacitors), or can include one or more systems for harnessing or generating electricity (such as fuel generators or solar panel arrays). Some disclosed embodiments provide weight reduction and space reduction of components in the aircraft, thereby improving the efficiency and performance of the aircraft. Considering the concern for safety in passenger transportation, the disclosed embodiments implement new and improved safety protocols and system redundancies in the event of a failure to minimize any single point of failure in the aircraft propulsion system. Some disclosed embodiments also provide new and improved methods for meeting aviation and transportation laws and regulations.

[0021] Figure 1A An exemplary eVTOL aircraft consistent with embodiments of the present disclosure is illustrated. As Figure 1A shown, in some embodiments, the distributed electric propulsion system of the eVTOL aircraft 100 can include twelve electric engines 110, which can be mounted on booms in front of and behind the main wing of the aircraft 100. The front electric engines 110 can tilt during flight between a horizontal orientation position (e.g., to generate forward thrust) and a vertical orientation position (e.g., to generate vertical lift). In terms of the propeller rotation direction, the front electric engines 110 can be clockwise type or counterclockwise type. The rear electric engines 110 can be fixed in a vertical orientation position (e.g., to generate vertical lift), and can also be clockwise type or counterclockwise type depending on the direction of propeller rotation.

[0022] The aircraft 100 can have various combinations of front and rear electric engines 110. For example, in some embodiments, the aircraft 100 can have six front electric engines 110 and six rear electric engines 110. In some other embodiments, the aircraft 100 can include four front electric engines 110 and four rear electric engines 110, or any other combination of front and rear engines 110. In some other embodiments, the number of front and rear electric engines is not equal.

[0023] In some embodiments, for vertical takeoff and landing (VTOL) missions, the front electric engine 110 and the rear electric engine 110 can provide vertical thrust during takeoff and landing. During the flight phase when the aircraft 100 is in the forward flight mode, the front electric engine 110 can provide horizontal thrust, and the propellers of the rear electric engine 110 can be stowed in a fixed position to minimize drag. The rear electric engine 110 can be actively stowed through position monitoring.

[0024] In some embodiments, in a conventional takeoff and landing (CTOL) mission, the front electric engine 110 can provide horizontal thrust for wing-borne takeoff, cruise, and landing. In some embodiments, during a CTOL mission, the rear electric engine 110 may not be used to generate thrust, and the rear propellers can be stowed in place.

[0025] The transition from vertical flight to forward flight and vice versa can be achieved via a tilt-propeller subsystem. The tilt-propeller subsystem can redirect thrust between the primary vertical direction during vertical flight mode and the primary horizontal direction during forward flight mode. A variable pitch mechanism can change the total blade pitch angle of the propeller-hub assembly of the front electric engine operating during the hover phase, transition phase, and cruise phase.

[0026] The tilt-propeller system can include a linear or rotary actuator to change the orientation of the propulsion system during operation. In some embodiments, the pitch of the propulsion system can be changed according to the orientation of the propulsion system. In some embodiments, the rotary actuator can include a motor, an inverter, and a gearbox. In some embodiments, the gearbox can include various types of gear interfaces to provide a gear reduction device capable of orienting the propulsion system. In some embodiments, the tilt-propeller system can include a redundant configuration such that multiple motors, inverters, and gearboxes are present and are gear-coupled. In some embodiments, a configuration using multiple motors, gearboxes, and inverters can allow a failed portion of the redundant configuration to be driven by the motors, inverters, and gearboxes of another portion of the configuration. In some embodiments, the gearbox configuration can also allow the tilt-propeller system to maintain the orientation of the propulsion system with or without the help of additional power provided by the system.

[0027] In some embodiments, the electric engine 110 can be housed on or attached to the pylon of the aircraft 100 and includes an electric motor, an inverter, and a gearbox. In some embodiments, the electric motor, the inverter, and the gearbox can be joined such that they share a central axis. In some embodiments, the torque originating from the electric motor can be transmitted from the propeller of the propulsion system to the gearbox. In some embodiments, the gearbox can provide a gear reduction and then send the torque back to the propeller via a main shaft through a bearing located inside the electric motor. In some embodiments, the inverter can be mounted on the rear of the gearbox such that the main shaft does not need to pass through the inverter when outputting torque to the propeller.

[0028] As Figure 1A shown, the aircraft 100 can be configured with a distributed electric propulsion system to enable vertical flight, forward flight, and transitional flight. The first six electric engines 110 (numbered 1 - 6 from left to right) have variable pitch propellers and achieve vertical takeoff and landing, transitional flight, and full wing-borne flight by tilting. The last six electric engines 110 (numbered 7 - 12 from left to right) are equipped with fixed pitch propellers, which operate during vertical takeoff and landing and transitional flight and are stowed in the minimum drag position for conventional flight. Flight control is an integrated fly-by-wire system with envelope protection and structural load limiting functions. The aircraft 100 will be equipped with advanced cockpit avionics, a flight management system, and sensors required to support the intended operations and system functions.

[0029] In some embodiments, an electric propulsion system (EPS) as described herein can generate thrust by supplying high voltage (HV) power to the electric engine 110, which in turn converts the HV power into mechanical shaft power for rotating the propeller. As described above, the aircraft 100 as described herein can have multiple electric engines 110 that are mounted on pylons in front of and behind the wings. The magnitude of the thrust generated by each electric engine 110 can be controlled by torque commands sent by a flight control system (FCS) to each electric engine 110 via a digital communication interface. Embodiments can include front electric engines 110 that are capable of changing their orientation or tilt. Additional embodiments include front engines that can be clockwise (CW) type or counterclockwise (CCW) type. The front electric engine propulsion subsystem can consist of a multi-blade variable pitch propeller and a variable pitch subsystem.

[0030] In some embodiments, the aircraft 100 includes a high voltage power supply (HVPS) system to supply high voltage (HV) power. The HVPS system is the power source on the aircraft 100 and is configured to distribute the stored electrical energy to other systems on the aircraft 100, including an electric propulsion system (EPS) that converts the electrical energy into mechanical rotational shaft power to generate thrust. As Figure 1AAs shown, the HVPS system of the aircraft 100 may include six battery packs 120 (numbered 1-6 from left to right), which are installed in the battery compartments of the wings of the aircraft 100. In some embodiments, the six battery packs 120 may have the same design to simplify design, manufacturing, and logistics. The battery packs 120 may supply power to one or more electric engines 110. Although six battery packs 120 are shown, the aircraft 100 may have any number of battery packs 120.

[0031] In some embodiments, a single battery pack 120 may be electrically connected to and supply power to multiple electric engines 110. For example, in some embodiments, the battery pack 120 may supply power to the electric engines 110 on either side of the longitudinal axis. In some embodiments, the battery pack 120 may supply power to the electric engines 110 on either side of the horizontal axis. In some embodiments, as Figure 1A shown, the battery pack 120 may supply power to two diagonally opposite electric engines 110. For example, battery pack 1 may supply power to electric engines 1 and 12. Battery pack 2 may supply power to electric engines 5 and 8. Battery pack 3 may supply power to electric engines 3 and 10. Battery pack 4 may supply power to electric engines 4 and 9. Battery pack 5 may supply power to electric engines 2 and 11. Battery pack 6 may supply power to electric engines 6 and 7. Thus, in the event of a loss of a battery pack 120, the impact on the roll or pitch moment can be reduced because the lift loss is balanced. In some embodiments, the battery packs 120 may supply power to different arrangements of the electric engines 110 to reduce the roll moment, pitch moment, or yaw moment that may be caused by the loss of a battery pack 120. For example, in some embodiments, the battery packs 120 may be connected to the electric engines 110 in any manner that balances lift and / or thrust with respect to the longitudinal axis and the horizontal axis of the aircraft.

[0032] Furthermore, the HVPS system includes a cross-link 130 having at least one fuse, allowing two or more battery packs 120 to be paired. Through the cross-link, the power for the electric engines 110 can be shared between the paired battery packs 120. Thus, multiple battery packs 120 can supply power to multiple electric engines 110 simultaneously. This arrangement provides redundancy and avoids single points of failure because each paired battery 120 can act as a backup battery for the other batteries. In the event of a failure of a battery pack 120, one or more connected battery packs 120 can continue to supply power to the connected electric engines 110 of the failed battery pack.

[0033] In some embodiments, as Figure 1AAs shown, a pair of battery packs 120 may include two battery packs 120. In some embodiments, a pair of two battery packs 120 may power a total of four electric engines 110. For example, battery pack 1 that powers electric engines 1 and 12 may be cross-connected with battery pack 4 that powers electric engines 4 and 9. Battery pack 2 that powers electric engines 5 and 8 may be cross-connected with battery pack 5 that powers electric engines 2 and 11. Battery pack 3 that powers electric engines 3 and 10 may be cross-connected with battery pack 6 that powers electric engines 6 and 7.

[0034] Figure 1B Another exemplary eVTOL aircraft consistent with embodiments of the present disclosure is illustrated. In some embodiments, electric engine 110 may include multiple motor stages, each motor stage independently powered by a different battery pack 120, such that if one battery pack 120 fails, only a portion of the EPU is not powered, and the EPU may continue to operate at a reduced power level. In some embodiments, electric engine 110 may include two partial motors. For example, battery pack 1 may power the first partial motors on electric engines 1, 6, 7, and 12. Battery pack 6 may power the second partial motors on electric engines 1, 6, 7, and 12. In some embodiments, different configurations may be used. For example, two battery packs may power partial motors on electric engines 1, 4, 9, and 12.

[0035] Figure 1C An electric engine 110 having two partial motors 191a and 191b consistent with embodiments of the present disclosure is illustrated. Partial motors 191a and 191b may be powered by different battery packs 120. The two partial motors 191a and 191b may operate independently to drive the blades of the EPU and may operate simultaneously to drive the blades at a higher power. Partial motors 191a and 191b are respectively driven by their own motor controllers 192a and 192b. In some embodiments, the power of the partial motors may be electrically separated such that each electric engine 110 has electrically separated backup power.

[0036] The above configuration is provided as an example, but different numbers and configurations of battery packs 120, electric engines 110, connections from the battery packs to the electric engines, and combinations of battery pack cross-links may be used. In some embodiments, each battery pack 120 may power a separate electric engine 110. For example, an aircraft may have four, six, eight, ten, twelve, or any number of electric engines 110, and the number of battery packs 120 may match the number of electric engines. In some embodiments, each battery pack 120 may power only one electric engine 110 and may be electrically isolated from all other battery packs 120. In some embodiments, each battery pack 120 may power one or more partial motors, and each electric engine may include two or more partial motors. Thus, each electric engine 110 may have a backup power source, but the battery packs 120 remain electrically isolated.

[0037] In some embodiments, each battery pack 120 may power multiple electric engines 110. As described above, the battery pack 120 may power multiple sets of electric engines 110 that are symmetric about one or more symmetry axes. In some embodiments, the battery pack 120 may power electric engines 110 that are symmetric about the longitudinal axis, the lateral axis, or both of the aircraft. For example, as described above, in some embodiments, different battery packs 120 may power diagonally symmetric electric engines 1 and 12, 2 and 11, 3 and 10, 4 and 9, 5 and 8, and 6 and 7.

[0038] In some embodiments, the battery pack 120 may power more than two electric engines 110. In some embodiments, the battery pack 120 may power two or more sets of diagonally symmetric electric engines. For example, in some embodiments, the battery pack 120 may power electric engines 3, 6, 7, and 10, where electric engines 3 and 10 are diagonally symmetric and electric engines 6 and 7 are diagonally symmetric. In some embodiments, the set of electric engines 110 powered by the battery pack 120 may include an inner pair of diagonally symmetric electric engines 110 and an outer pair of diagonally symmetric electric engines 110.

[0039] In some embodiments, the battery pack 120 may power four or more electric engines 110 configured to be symmetric about a longitudinal symmetry axis. For example, the battery pack 120 may power electric engines 1, 6, 7, and 12. In some embodiments, in each of the above configurations, the battery pack 120 may supply power to one or more partial motors, and each electric engine 110 may include two or more partial motors. Thus, each electric engine 110 may have a backup power source, but the battery packs 120 remain electrically isolated.

[0040] In some embodiments, some or all of the battery packs 120 in the battery pack are interconnected. As described above, the cross-links 130 can allow each battery pack 120 to act as a backup power source for another battery pack. For example, in some embodiments, the first battery pack can directly power a first number of electric engines, and the second battery pack 120 can directly power a second number of electric engines. The first battery pack and the second battery pack 120 can be cross-connected together to form a battery pack unit. Thus, each battery pack in the unit can serve as a backup battery pack for the other. When a battery pack in the unit fails, the failed battery pack can be disconnected, and the electric engine 110 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.

[0041] As described above, in some embodiments, the battery pack unit can include two battery packs 120, where each battery pack 120 powers a certain number of electric engines 110. As described above, in some embodiments, each battery pack 120 can power two diagonally symmetric electric engines 110. Thus, each battery pack unit can power a total of four electric engines 110, and each electric engine has a backup battery pack. In some embodiments, each battery pack 120 in the battery pack unit can power four electric engines 110 (including two sets of diagonally symmetric electric engines 110). Thus, each battery pack unit can power a total of eight electric engines 110, and each electric engine has a backup battery pack.

[0042] In some embodiments, the battery pack unit can include three battery packs 120, where each battery pack powers a certain number of electric engines 110. For example, in some embodiments, each battery pack 120 can power two diagonally symmetric electric engines 110. Thus, each battery pack unit can power a total of six electric engines 110, and each electric engine 110 has two backup battery packs. In some embodiments, each battery pack 120 in the battery pack unit can power four electric engines 110 (including two sets of diagonally symmetric electric engines 110). Thus, each battery pack unit can power a total of twelve electric engines 110, and each electric engine has two backup battery packs.

[0043] In some embodiments, the battery pack unit may include four battery packs 120, where each battery pack powers a certain number of electric motors. For example, in some embodiments, each battery pack may power two diagonally symmetric electric motors 110. Thus, each battery pack unit can power a total of eight electric motors 110, and each electric motor has three backup battery packs. In other embodiments, each battery pack 120 in the battery pack unit may power four electric motors 110 (including two sets of diagonally symmetric electric motors 110). Thus, each battery pack unit can power sixteen electric motors 110, and each electric motor has three backup battery packs.

[0044] In some embodiments, the electric motor 110 includes a single motor powered by one or more battery packs 120. In some embodiments, each electric motor 110 may include two or more partial motors, and the battery pack 120 may power the partial motors. In some embodiments, the electric motor 110 powered by the above configuration may include powering the partial motors of the battery pack. For example, in some embodiments, each electric motor 110 may include two partial motors, and the battery pack unit may power the partial motors of electric motors 2, 4, 6, 7, 9, and 11. The second battery pack unit may power the partial motors of electric motors 2, 3, 6, 7, 10, and 11. The third battery pack unit may power the partial motors of electric motors 1, 3, 5, 8, 10, and 12. The fourth battery pack unit may power the partial motors of electric motors 1, 4, 5, 8, 9, and 12. Thus, each electric motor 110 will receive backup power through another partial motor. As described above, each battery pack unit may include one or more battery packs. For example, the battery pack unit may include one, two, three, or four battery packs.

[0045] In some embodiments, each electric motor may include two partial motors, and the battery pack unit may power the partial motors of electric motors 2, 3, 4, 5, 8, 9, 10, and 11. The second battery pack unit may power the partial motors of electric motors 1, 2, 5, 6, 7, 8, 11, and 12. The third battery pack unit may power the partial motors of electric motors 1, 3, 4, 6, 7, 9, 10, and 12. Thus, each electric motor 110 will receive backup power through another partial motor. As described above, each battery pack unit may include one or more battery packs. For example, the battery pack unit may include one, two, three, or four battery packs.

[0046] In some embodiments, each electric engine may include two partial motors, and the battery pack unit may power the partial motors of electric engines 3, 4, 9, and 10. The second battery pack unit may power the partial motors of electric engines 3, 4, 9, and 10. The third battery pack unit may power the partial motors of electric engines 2, 5, 8, and 11. The fourth battery pack unit may power the partial motors of electric engines 2, 5, 8, and 11. The fifth battery pack unit may power the partial motors of electric engines 1, 6, 7, and 12. The sixth battery pack unit may power the partial motors of electric engines 1, 6, 7, and 12. Thus, each electric engine 110 will receive backup power through another partial motor. As described above, each battery pack unit may include one or more battery packs. For example, the battery pack unit may include one, two, three, or four battery packs. Different configurations of the battery pack 120, the electric engine 110, the connection from the battery pack to the electric engine, and the battery pack cross-link may be selected to optimally balance the aircraft power demand, system redundancy, and fault tolerance.

[0047] Figure 1D FIG. illustrates a schematic diagram of a high-voltage power system for an eVTOL aircraft consistent with an embodiment of the present disclosure. As Figure 1B shown, the eVTOL aircraft may include a battery assembly that includes electrically separated battery pack units (such as 160, 162, and 164). As described above, each battery pack unit may include battery packs 120 cross-connected together. In some embodiments, the battery pack unit may include battery packs 120 that ensure the controllability of the aircraft is maintained in the event of the loss of a battery pack unit. Thus, the aircraft can still be controlled in the event of the loss of a battery pack unit. As described above, in some embodiments, the battery pack unit may include battery packs 120 that power the electric engine 110 on opposite sides of one or more symmetry axes. Thus, in the event of the loss of a battery pack unit, the impact on roll, pitch, or yaw moments can be reduced since the loss of lift and / or thrust is balanced. In some embodiments, the power loss or power reduction caused by the failure of a battery pack unit will have a substantially symmetric impact on the roll, pitch, and / or yaw of the aircraft (e.g., <±5%, <±10%, <±15%, <±20%, or <±25% asymmetry). In some embodiments, the battery pack unit may include battery packs 120 to reduce the total amount of high-voltage lines between the battery packs. In some embodiments, the battery pack unit may include battery packs 120 to minimize the power demand.

[0048] In some embodiments, as Figure 1DAs shown, the HVPS system may include three electrically isolated battery pack units. For example, in some embodiments, battery pack unit 160 may include battery packs 1 and 4 that power electric engines 1, 4, 9, and 12. Battery pack unit 162 may include battery packs 2 and 5 that power electric engines 2, 5, 8, and 11. Battery pack unit 164 may include battery packs 3 and 6 that power electric engines 3, 6, 7, and 10. Thus, each battery pack unit may include two paired battery packs 120 that simultaneously power four electric engines 110. When one of the battery packs 120 in a battery pack unit fails, the other paired battery pack 120 will continue to power the four electric engines.

[0049] In some embodiments, each battery pack unit 160, 162, 164 may include a high-voltage bus to cross-connect the battery packs 120 within the battery pack unit. In some embodiments, cross-link 130 connects two high-voltage channels, each channel feeding one or more electric engines 110. For example, in some embodiments, cross-link 130 may be connected to the high-voltage channels of each battery pack before the channels split to power multiple electric engines 110 (e.g., to power two electric engines). The cross-link may further include a bus that connects the negative voltage channels after the negative voltage channels are combined (e.g., after powering two electric engines).

[0050] In some embodiments, each cross-link 130 may include at least one fuse to disconnect the cross-link in the event of a cross-link failure. For example, fuses 131, 132, and 134 may be located on the cross-links of the positive high-voltage channels in battery pack units 160, 162, and 164. In some embodiments, the fuses may be pyrotechnic fuses. As described in further detail below, the battery management system of the connected battery packs 120 may determine a failure in the cross-link, such as a short circuit or overcurrent condition, and blow the associated pyrotechnic fuse. Thus, the cross-link may be disconnected, and further damage to HVPS system components (e.g., electric engines, batteries, EPUs) may be avoided. Further, the electric engines 110 will still receive power from the paired battery packs 120 in the battery pack unit. For example, when a cross-link fails, pyrotechnic fuse 131 may be blown, but electric engines 1 and 12 will still receive power from battery pack 1, and electric engines 4 and 9 will still receive power from battery pack 4.

[0051] In some embodiments, additional pyrotechnic fuses may be present on the cross-link connections of the negative high voltage channels. For example, pyrotechnic fuses 170, 172, and 174 may be located on the cross-links in battery pack units 160, 162, and 164, respectively. This configuration can provide additional redundancy for the system. If the fuses on the positive cross-link connections fail, the fuses on the negative cross-link connections can act as backup fuses and vice versa. For example, in some embodiments, if the fuses on the positive cross-link connections do not blow after being commanded to blow, the connected battery management system can instruct the negative cross-link fuses to blow. Further, in some embodiments, each positive cross-link may have two fuses controlled by two associated battery packs, and each negative cross-link may have two fuses controlled by two associated battery packs.

[0052] In some embodiments, the HVPS system may include load disconnect devices to disconnect a portion of the HVPS circuit in the event of a failure (e.g., a short circuit or overcurrent condition) in a downstream electric engine, downstream EPU, or other downstream distribution circuit. In some embodiments, the load disconnect devices may be located directly upstream of the electric engines. For example, in some embodiments, load disconnect devices 109, 111, 112, and 113 may be located on the high voltage channels supplying power to engines 1, 12, 4, and 9, respectively. Load disconnect devices 114, 115, 116, and 117 may be located on the high voltage channels supplying power to engines 2, 11, 5, and 8, respectively. Load disconnect devices 118, 119, 121, and 122 may be located on the high voltage channels supplying power to engines 3, 10, 6, and 7, respectively.

[0053] In some embodiments, the load disconnect device is a pyrotechnic fuse. In the event of a failure of a downstream component, the pyrotechnic fuse may receive a signal (e.g., from the battery management system of the connected battery) and blow the fuse. Thus, the downstream component can be disconnected, and further damage to other devices (such as electric engines, batteries, EPUs) can be avoided. Further, the remaining electric engines 110 in the battery pack unit will still receive power from the connected battery pack 120. For example, when a device or line downstream of pyrotechnic fuse 109 fails, pyrotechnic fuse 109 may blow, but electric engines 12, 4, and 9 will still receive power from battery packs 1 and 4. Further, in some embodiments, the load disconnect device may include a contactor, and the battery management system can command the contactor to disconnect the circuit. In some embodiments, both the contactor and the fuse can be used to provide additional redundancy, and the pyrotechnic fuse can serve as a backup fuse for the contactor.

[0054] In some embodiments, the HVPS system may include a high-voltage charging channel that allows all battery packs 120 to be charged from the same charging port. The high-voltage charging channel may include a charging disconnect device. In some embodiments, the charging disconnect device may be positioned downstream of the common charging bus on the positive charging side. For example, disconnect devices 140, 142, 144, 146, 148, and 150 may disconnect battery packs 1, 4, 5, 2, 3, and 6, respectively. Similarly, in some embodiments, additional charging disconnect devices may be positioned upstream of the common charging bus on the negative charging side. For example, disconnect devices 141, 143, 145, 147, 149, and 151 may disconnect battery packs 1, 4, 5, 2, 3, and 6, respectively.

[0055] In some embodiments, the charging disconnect device is a contactor, such as Figure 2A K4 Pos and K4 Neg in. The charging contactor may act as a redundancy measure to disconnect the charging of battery pack 120. As described in further detail below, battery pack 120 may report a charging problem to the charging control unit (CCU). For example, battery pack 120 may report a short circuit or overcurrent condition in battery pack 120 or in the high-voltage charging channel. In some embodiments, if the CCU fails to stop charging, battery pack 120 may command the charging contactor to disconnect the charging channel. In some embodiments, battery pack 120 may automatically command the charging contactor to disconnect the charging channel without waiting for the CCU to fail. In some embodiments, after commanding the CCU to stop charging and / or disconnect the battery pack 120 in which a charging problem is detected, battery pack 120 and / or the CCU may command other battery packs 120 to disconnect from the charging channel. By disconnecting battery pack 120 when a charging problem is detected, damage to HVPS components can be avoided.

[0056] Figure 2A Illustrated is a circuit diagram of a high-voltage junction box (HVJB) consistent with embodiments of the present disclosure. The HVJB 222 may be electrically connected to the HV load 210 to provide high-voltage power. Specifically, a DC / DC converter and energy storage element BT1 (e.g., battery cells connected in parallel and in series) in the battery management system (BMS) may be used to provide high-voltage power. The DC / DC converter and energy storage element BT1 are each connected to each of the HV loads through a combination of a pre-charge resistor (e.g., resistor R1) or current sense resistors (e.g., resistors R2 - R6), switching devices K1 - K5 (e.g., HV contactors, relays, and / or controllers), and active and passive fuses (e.g., F1 - F7) to guard against various fault conditions (such as overcurrent, short circuit, etc.). In some embodiments, fuses F1 - F7 may be as referred to above Figure 1BOne or more fuses in the detailed fuses. For example, in some embodiments, fuses F2 EE1, F3 EE2, and F4 Xlink may correspond to Figure 1D the fuses 109, 111, and 131 detailed in

[0057] Fuse F1 may be a battery pack fuse that disconnects a failing battery pack 120 from the rest of the HVPS system. In some embodiments, F1 may be a pyrotechnic fuse. When the battery pack 120 fails, the pyrotechnic fuse F1 may receive a signal (e.g., from an associated battery management system) and blow the fuse F1. Thus, further damage to other devices (e.g., electric engine, EPU, connected battery packs) can be avoided. Further, the electric engine 110 will still receive power from the paired battery pack 120 within the battery pack unit. For example, when a battery pack fails, the pyrotechnic fuse F1 of battery pack 1 may be blown, but the electric engines 1, 12, 4, and 9 can still receive power from battery pack 4.

[0058] The circuit arrangement in the high voltage junction box (HVJB) 222 provides charging flexibility by allowing the auxiliary load and / or the electric engine and actuator to be powered on or off during charging. For example, the battery pack 120 can be charged while the remaining HVPS circuit remains disconnected. The charging contactors K4 positive and K4 negative can be closed to allow the battery pack 120 to charge. At the same time, the main contactors K1 and K2 and the pre-charge contactors (and / or relays) K3 and K5 can be opened to prevent powering on the remaining HVPS circuit. Further, the battery pack 120 can be charged while the auxiliary load is connected but the electric engine and actuator remain disconnected. The charging contactors K4 positive and K4 negative can be closed to allow the battery pack 120 to charge. At the same time, after the pre-charge contactor (and / or relay) K3 completes the pre-charge of the auxiliary load, the main contactors K1 and K2 can be closed, and K4 EE can remain open. Further, the battery pack 120 can be charged while all loads are connected. The charging contactors K4 positive and K4 negative can be closed to allow the battery pack 120 to charge. At the same time, after the pre-charge contactors (and / or relays) K3 and K5 complete the pre-charge of the connected loads, the main contactors K1 and K2 can be closed, and K4 EE can be closed.

[0059] In some embodiments, the input device may allow a person to select a charging mode for the aircraft. For example, the person requests one of the above three charging modes for charging through the input device. In some embodiments, the input device may be a physical switch, button, and / or joystick. In some embodiments, the input device may be a user interface element provided on a display screen or a control panel. In some embodiments, the input device may be a processor that can receive voice commands for manually selecting and / or requesting a mode switch. The input device may include any device that allows a person to select a desired charging mode. In some embodiments, the input information is transmitted to the BMS 270, and the BMS 270 may control the contactor according to the requested charging mode.

[0060] Figure 2B FIG. illustrates a schematic diagram of a high voltage junction box 222 (HVJB) consistent with an embodiment of the present disclosure. In some embodiments, each battery pack 120 includes a HV power distribution unit 211, a battery management system (BMS 270), and a pyrotechnic fuse redundant trigger board (PRT 280) housed within the HVJB 122. Each unit may be a hardware device such as a computer, a processor, or a microprocessor. The BMS 270 may be configured to monitor voltage, temperature, current, and isolation resistance. The BMS 270 may control the battery pack contactor and the pyrotechnic fuse to prevent a fault condition. As described in further detail below, the BMS 270 may communicate with various systems inside and outside the HVJB 222. The BMS 270 may include a battery management unit (BMU 271) that can receive voltage, current, resistance, and temperature sensing signals from the cell stack assembly 224 and / or the HV power distribution unit 211. The BMS 270 may further include a cell management unit (CMU) 272 to monitor the voltage of each group of 7 parallel cells (i.e., 1S-7P cell group) connected in series in a 14S-7P cell block. The CMU may also be used to monitor the temperature of the 14S-7P cell block. The CMU 272 obtains the measurement values of all cell groups in the battery pack 120 and transmits these measurement values to the BMU 271.

[0061] The BMU 271 can monitor the output current of each connection load in the connection load. The BMU 271 can be internally powered by the battery cell stack assembly 224 and continuously monitor the battery status even when the battery is not installed in the aircraft 100. By monitoring the battery pack 120, cell block, and cell group parameters, the BMU can prevent conditions that adversely affect safety or performance, such as overvoltage, undervoltage, overheating, underheating, electrical insulation loss, short circuit, overcurrent, etc. The diagnostic function of the BMU 271 allows for fault detection and isolation through built-in test (BIT). Additionally, the BMU 271 performs calculations of the state of charge (SOC), state of health (SOH), fault condition (e.g., short circuit or overcurrent), power state (SOP), energy state (SOE), and temperature state (SOT) of the battery pack 120. The BMU 271 also controls and monitors bus pre-charging, provides fuse and contactor commands, and communicates with various systems inside and outside the HVJB 222.

[0062] The HV power distribution unit 211 in the HVJB 222 can include an HV contactor 212 and a combination of active and passive fuses (e.g., pyrotechnic fuse 213 and fuse 214) to prevent overcurrent and short circuit conditions. In some embodiments, the contactor 212 can correspond to Figure 2A one or more of the switching devices K1-K7 (e.g., HV contactors) detailed in Figure 2A Similarly, the pyrotechnic fuse 213 and fuse 214 can correspond to one or more of the fuses F1-F8 detailed in

[0063] In some embodiments, the pyrotechnic fuse redundant trigger board (PRT 280) may be located within the HVJB 222. In other embodiments, the BMS 270 may communicate with the PRT 280 located outside the HVJB 222. The BMS 270 may detect a fault event and send a command signal to the PRT 280 to cause the corresponding pyrotechnic fuse driver to blow the fuse. For example, in some embodiments, the HV power distribution unit 211 may receive sensor signals from current sensors (e.g., resistors R3 - R6) and provide information about the condition (e.g., voltage, current, or temperature) of the connected load at a point in the HVPS system to the BMU 271. Based on the received information, the BMU 271 may determine a fault condition (e.g., because the value is outside a predetermined range) and send a command to the PRT 280 to blow the associated pyrotechnic fuse. Thus, the fault condition may be disconnected from the rest of the HVPS circuit, protecting the remaining devices and lines. In some embodiments, the BMU 271 may directly monitor the sensors instead of receiving information through the HV power distribution unit 211.

[0064] In some embodiments, the battery packs 120 may communicate with each other, for example, through the BMS 270. The battery packs 120 may use information about the state of one or more paired battery packs 120 in the battery pack unit to help determine whether an overcurrent condition has occurred. For example, the battery packs 120 may determine an expected operating range (e.g., voltage, current, etc.) based on the state of the battery pack and the communication state of the battery packs 120 within the battery pack unit. In some embodiments, the HVJB 122 may further provide a redundant active trigger board configured to enable the pyrotechnic fuse driver to activate one or more pyrotechnic fuses when the BMS 270 fails to enable the pyrotechnic fuse driver. See US11,710,957, which is incorporated herein by reference.

[0065] The control MCU (CCU 263) in the charging port assembly 262 may interface with an external battery charger and communicate with the BMU 271 on the six installed battery packs 120. The unit may be a hardware device such as a computer, a processor, or a microprocessor. In some embodiments, the CCU 263 may be a single PCBA with one microcontroller that manages the total power delivery to each battery pack 120 during charging. As shown in FIG. 2, the CCU 263 may perform a handshake between the ground charging subsystem 274 and the BMU 271 and may command the BMU 271 to open or close the contactor 212, such as Figure 2AThe contactors K6-K7 detailed in []. The CCU 263 can perform active detection and protection features for overvoltage protection. During the charging operation, the BMU 271 in each battery pack 120 can maintain full control and continuously monitor their battery pack 120.

[0066] Figure 3 FIG. illustrates a schematic diagram of a charging port assembly (CPA 262) consistent with an embodiment of the present disclosure. The charging port assembly 262 includes a charging port 330, thereby providing a communication connection via power line communication 335 and HV power transmission via the HV power channel 336. In some embodiments, the charging port 330 can be a type JI 772 charging port including various pins and connection points to allow connection to a ground service system (GSS 300) (e.g., via a plug). The charging port 330 can include one or more proximity pins to detect a high-voltage connection between the GSS 300 and the charging port 330. Upon detecting a connection to the GSS 300, the charging port 330 can engage a latch that prevents the high-voltage power 336 from disconnecting under a charging load. After charging is complete, the charging port 330 can automatically unlock the connection or enable manual unlocking.

[0067] The charging port assembly 262 can include a charging control unit (CCU 263) that communicates with the charging port 330, for example, via a communication line 333. The CCU 263 can further provide latch control 331, lighting change 334, and monitor and respond to the temperature of various components 332. The CCU 263 can monitor the temperature on the inlet side of the charging port 330. If the temperature becomes too high, the CCU 263 can command the ground service system 300 to abort charging. The CCU 263 receives status updates from the battery pack 120 and provides commands to the battery pack 120 to control their charging levels by opening and closing the battery pack charging contactors (e.g., Figure 2A K6-K7 in []). As described in detail above with reference to Figure 2B In some embodiments, the CCU 263 can communicate with the battery management system (BMS 270) of each battery pack, for example, via a battery management unit (BMU 271). The BMS 270 can send battery pack information to the CCU 263, including information about the battery pack connection status (e.g., whether the battery pack is connected to the HVPS system), charge state (SOC), health state (SOH), fault conditions (e.g., short circuit or overcurrent), power state (SOP), energy state (SOE), and temperature state (SOT).

[0068] The CCU 263 can provide commands to the BMS 270 to disconnect or close the battery pack charging contactor. In some embodiments, each battery pack 120 can have a separate low-voltage CAN communication line that connects the battery pack 120 to the CCU 263. In some embodiments, the CAN communication lines can be shared among one or more of the battery pack units. For example, HV battery packs 1 and 4 can communicate with the CCU 263 via CAN 351. HV battery packs 2 and 5 can communicate with the CCU 263 via CAN 352. HV battery packs 3 and 6 can communicate with the CCU 263 via CAN 353. As described in further detail below, the CCU 263 can make various power supply and cooling requests for the GSS 300 (e.g., via the charging port 330) based on information received from the battery packs 120.

[0069] The charge control unit 263 can determine the battery pack charging contactor commands based on various criteria. In some embodiments, the CCU 263 can determine the required battery pack charge level based on flight information. For example, in some embodiments, the CCU 263 can receive flight information from the GSS 300 via communication lines 335 and 333, for example. The GSS 300 can receive flight information via a wired or wireless connection to a computer, laptop, ipad, mobile device, or any other device capable of providing flight information. In some embodiments, the charging port assembly 262 can provide a direct wired or wireless connection to a computer, laptop, ipad, mobile device to receive flight information directly. In some embodiments, the CCU 263 can receive flight information from the flight control system 230 of the aircraft.

[0070] Flight information can include flight mission information, such as the location of the destination, the distance to the next destination, or the expected flight time required to reach the next destination. The flight mission information can include the expected type of flight. For example, the flight mission information can include the duration or distance to be covered in each flight mode. In some embodiments, the flight modes can include winged flight, thrust and lift assisted flight, thrust assisted flight, and lift assisted flight. In some embodiments, the flight mission information can include the expected EPU output during the entire flight, for example, as a power unit or as a percentage of the maximum EPU power. In some embodiments, flight mission information can be provided for each EPU on the aircraft.

[0071] Flight mission information can include information about predicted weather conditions throughout the flight. Weather conditions may include temperature, pressure, wind conditions, and expected rainfall throughout the flight. Flight mission information can include the expected weight of the aircraft, for example, based on the number of passengers or amount of cargo. The weight of the aircraft can be predicted or measured (e.g., if the aircraft is loading passengers or cargo).

[0072] Flight information can include historical battery information. For example, in some embodiments, the battery information can include historical battery consumption of each battery pack along a particular flight path. The battery information can further include details about flight mode, weight, and weather for the charge control unit 263 to determine its relevance to the upcoming flight mission.

[0073] Further, flight information for multiple subsequent flights can be received and analyzed. In some embodiments, if the aircraft will make multiple trips without the ability to recharge, flight information can be collected and analyzed for all subsequent flights to ensure that the aircraft has sufficient power for each trip. In some embodiments, the aircraft may have time for partial charging before subsequent trips. Thus, the flight information can include information about subsequent trips and information about the amount of recharge available between trips. By receiving this information, the CCU 263 can ensure that the battery pack 120 has sufficient power to support a sufficient portion of the subsequent trip. The CCU 263 can use the flight information to determine the required charge level for each battery pack 120.

[0074] The charge control unit (CCU 263) can determine the battery pack charging contactor commands based on the current state of each battery pack 120 received from the BMS 270, including the energy state and / or charge state of each battery pack 120. The CCU 263 can determine how much additional power is needed to meet the required charge level based on the current charge level of each battery pack. Further, in some embodiments, the CCU 263 can consider the battery pack configuration when charging the battery pack 120. The CCU 263 can determine to charge each battery pack 120 within the battery pack unit to the same charge level. Thus, the CCU 263 can charge all battery packs 120 within the battery pack unit to the highest charge level required by any battery pack 120 within that unit. As the battery pack 120 charges, the CCU 263 can continue to receive updates on the charge level of each battery pack and keep the battery pack charging contactor closed to allow charging until the required charge level is reached.

[0075] Further, the charge control unit (CCU 263) can determine a battery pack charging contactor command based on a fault condition, a state of health, or a temperature state received from the BMS 270. In some embodiments, the CCU 263 can disconnect the contactor of the battery pack 120 (prohibit charging) based on receiving information that the battery pack 120 has failed (e.g., experienced a short circuit or an overcurrent condition). Further, the CCU 263 can disconnect the contactor of the battery pack 120 (prohibit charging) based on the state of health of the battery pack dropping below a set level or based on the temperature of the battery pack exceeding a set level. The CCU 263 can continue to monitor the fault condition, the state of health, or the temperature state of the battery pack 120, and close the contactor (allow charging) when the condition is remedied.

[0076] The charge control unit 263 can send a cooling command to the GSS 300 based on the state of temperature information received from the battery pack 120, e.g., via communication lines 333 and 335. In some embodiments, the CCU 263 can send a desired battery pack temperature or a desired coolant flow rate. The ground charge subsystem 310 can communicate this information to the thermal regulation subsystem 320. The thermal regulation subsystem 320 can control one or more condensers and associated coolant control valves to meet the cooling requirements.

[0077] The charge control unit (CCU 263) can signal the state of the battery pack 120 to the charging service personnel throughout the charging process. In some embodiments, the CCU 263 can signal a problem (e.g., a battery pack fault, poor health, or excessive temperature) via the lighting line 334. For example, in some embodiments, a light can be turned on or change color to indicate a problem. Alternatively or additionally, the CCU 263 can communicate problem details (e.g., problem type, associated battery pack, etc.) to the ground service system 300 via communication lines 333 and 335. The ground service system 300 can provide these details via a display, computer, laptop, ipad, mobile device, or any other device capable of communicating the information to the charging service personnel.

[0078] The charge control unit (CCU 263) can determine that each battery pack 120 has reached the desired charge level and send a signal indicating charging completion to the ground service system 300. When determining that no charge is received from the GSS 300, the CCU 263 can provide a signal to the charging port, e.g., via latch control 331, to automatically unlock the connection to the GSS 300 or allow manual unlocking of the connection.

[0079] Figure 4The figure illustrates a flowchart for detecting an emergency transponder consistent with an embodiment of the present disclosure. In some embodiments, the process may be performed by each battery management system 270 of the battery pack 120. At step 401, the processor receives acceleration information. In some embodiments, the acceleration information may be received directly from a sensor (e.g., an accelerometer), while in other embodiments, the acceleration information may be received from a different processor (such as a processor associated with the flight control system of an aircraft). At step 402, the processor receives a high voltage interlock loop (HVIL) continuity status (e.g., from the battery management system (BMS270)) indicating whether the low voltage emergency cut-off loop has been cut. For example, the BMS270 may determine that the cut-off loop has been cut based on detecting a loss of current. The information collected in steps 401 and 402 may be received sequentially or simultaneously. Further, in some embodiments, the collected information may include a timestamp indicating when it was collected. In other embodiments, the processor may assign a time based on the time at which it received the information.

[0080] At step 403, the processor may determine whether the emergency transponder has performed a cut-off of the low voltage emergency cut-off loop. The processor may make this determination based on the acceleration information and the HVIL continuity status. If the acceleration information indicates a collision (e.g., exceeding a threshold) at an earlier time than the HVIL continuity status indicates the cut-off loop, it is determined that the emergency transponder has cut the low voltage emergency cut-off loop. However, if the HVIL continuity status indicates the cut-off loop at an earlier time than the acceleration information indicates a collision, no emergency response is detected. Further, if the acceleration information does not indicate a collision or the HVIL continuity status does not indicate a cut, no emergency response is detected. At step 404, if it is determined that the emergency transponder has performed a cut-off, the processor may send a command to blow one or more battery pack fuses to de-energize at least a portion of the high voltage power system. In some embodiments, the processor may determine which battery pack 120 to blow based on which battery pack 120 is associated with the cut-off loop. For example, in some embodiments, cut-off loop 1 may be connected to battery pack 1. The processor may determine that the emergency transponder has cut off cut-off loop 1, and the processor may instruct battery pack 1 to blow the pyrotechnic fuse of battery pack 1, such as Figure 2A the fuse F1 in. In some embodiments, based on determining that the emergency transponder has cut off any loop in the loop, the processor may blow the pyrotechnic fuses associated with the battery pack 120 and any connected battery packs 120. For example, referring to Figure 1A, based on determining that the emergency transponder has cut off Loop 1 associated with Battery Pack 1, the processor may fuse the pyrotechnic fuses associated with Battery Packs 1 and 4. In some embodiments, based on determining that the emergency transponder has cut off any loop in the loop, the processor may fuse the pyrotechnic fuses associated with all Battery Packs 120.

[0081] At step 405, the processor may determine whether the collision detection is false. The processor may determine that the acceleration information indicates a collision, but the HVIL continuity status indicates that no loop has been cut off. Further, the processor may collect or obtain information on whether the flight control system is in the ground mode. If the processor determines that the aircraft flight control system is in the ground mode, the processor may determine that the collision detection is false. However, if the processor determines that the flight control system is not in the ground mode (e.g., in the flight mode), a false collision will not be determined. In some embodiments, the "ground mode" may be a mode selected by the pilot through an interface when operating the aircraft on the ground.

[0082] At step 406, if it is determined that the collision detection is false, Condition 1 will be reset to indicate that no collision has been detected, and the processor will re-collect the acceleration information. At step 407, if it is not determined that the collision detection is false, Condition 1 will not be reset, and the processor will continue to monitor whether the HVIL continuity status indicates a cut-off loop in Condition 2 of step 403.

[0083] Figure 5A A plan view showing the routing of the cut-off loop through the tail of an eVTOL aircraft in accordance with an embodiment of the present disclosure is illustrated. As described above, each cut-off loop may be connected to a single Battery Pack 120. Thus, 6 cut-off loops may be routed from the Battery Pack 120 located in the wing or elsewhere to the tail of the aircraft. This routing ensures that the cut-off loop is easily cut at the tail of the aircraft away from the high-voltage power system that extends towards the front of the aircraft between the battery, the electric engine, and other aircraft equipment. The first transponder may cut off one or more loops to power off the Battery Pack 120 without the risk of cutting off the energized high-voltage power line, thereby increasing safety. In some embodiments, each cut-off loop may be routed separately. In some embodiments, the cut-off loop may be routed together with one or more Battery Packs 120 (e.g., in a bundle). For example, the cut-off loops associated with the connected battery packs may be bundled together, or the cut-off loops associated with the wing of the aircraft may be bundled together. In some embodiments, the cut-off loops of the Battery Pack 120 may all be routed in a single bundle.

[0084] Figure 5BIllustrated is a cross-sectional schematic diagram of a cut-off circuit line arranged at the tail of an eVTOL aircraft in accordance with an embodiment of the present disclosure. As described above, the cut-off circuit is arranged at the tail of the aircraft to increase the safety of the emergency transponder. Additionally, the cut-off circuits can be arranged in a manner that allows them to be easily accessed by the first transponder. For example, in some embodiments, the cut-off circuits can be arranged towards the periphery of the aircraft so that they are easier to detect and cut. The cut-off circuits can be color-coded and include descriptive tags at set intervals to ensure that the first transponder can easily identify them.

[0085] The embodiments can be further described using the following clauses:

[0086] Clauses A to C for aircraft charging:

[0087] Clause set A: A charging system for an aircraft, the charging system comprising: a plurality of electric propulsion units (EPUs); a plurality of battery packs configured to supply power to the plurality of EPUs; a charging control unit configured to: determine a target charge level for each of the plurality of battery packs; receive charge status information from each of the plurality of battery packs; and command one of the plurality of battery packs to disconnect charging when it is determined that the target charge level of the battery pack has been reached.

[0088] 2. The system according to clause A1, wherein the charging control unit is further configured to: command the ground charging subsystem to stop charging when it is determined that the target charge level of each of the plurality of battery packs has been reached.

[0089] 3. The system according to clause A1 or A2, wherein the charging control unit determines the target charge level for each of the plurality of battery packs based on flight information.

[0090] 4. The system according to clause A3, wherein the flight information includes at least one of the following: distance to the next destination, flight time to the next destination, flight mode to the next destination, expected weather conditions, historical battery consumption information, or charging availability information.

[0091] 5. The system according to clause A4, wherein determining the target charge level for at least one of the plurality of battery packs is different from the target charge level for another of the plurality of battery packs.

[0092] 6. The system according to any one of clauses A1 to A5, wherein the charging control unit is further configured to: receive fault status information from each of the plurality of battery packs; and command one of the plurality of battery packs to disconnect from charging when it is determined that a fault has occurred in the battery pack.

[0093] 7. The system according to clause A6, wherein the charging control unit is further configured to: command the ground charging subsystem to stop charging when it is determined that a fault has occurred in the battery pack.

[0094] 8. The system according to clause A7, wherein the charging control unit is further configured to: provide an indication to the ground charging subsystem specifying which one of the plurality of battery packs has a fault.

[0095] 9. The system according to any one of clauses A6 to A8, wherein the fault status information includes at least one of overcurrent information, short - circuit information, battery pack health status information, or battery pack temperature information.

[0096] 10. The system according to any one of clauses A1 to A9, wherein the charging control unit is further configured to: receive temperature status information from each of the plurality of battery packs; and provide a cooling command to the ground subsystem based on the temperature status information.

[0097] Clause set B: A control unit for charging an aircraft, the control unit comprising: a charging control unit configured to: determine a target charge level for each of the plurality of battery packs; receive charge status information from each of the plurality of battery packs; and command one of the plurality of battery packs to disconnect from charging when it is determined that the target charge level of the battery pack has been reached.

[0098] 2. The control unit according to clause B1, wherein the charging control unit is further configured to: command the ground charging sub - control unit to stop charging the plurality of battery packs when it is determined that the target charge level of each of the plurality of battery packs has been reached.

[0099] 3. The control unit according to clause B1 or B2, wherein the charging control unit determines the target charge level for each of the plurality of battery packs based on flight information.

[0100] 4. The control unit according to clause B3, wherein the flight information includes at least one of the following: distance to the next destination, flight time to the next destination, flight mode to the next destination, expected weather conditions, historical battery consumption information, or charge availability information.

[0101] 5. The control unit according to clause B4, wherein the determined target charge level of at least one battery pack among the plurality of battery packs is different from the target charge level of another battery pack among the plurality of battery packs.

[0102] 6. The control unit according to any one of clauses B1 to B5, wherein the charge control unit is further configured to: receive fault status information from each of the plurality of battery packs; and command one of the plurality of battery packs to disconnect from charging when it is determined that the battery pack has a fault.

[0103] 7. The control unit according to clause B6, wherein the charge control unit is further configured to: command the ground charge sub-control unit to stop charging the battery pack when it is determined that the battery pack has a fault.

[0104] 8. The control unit according to clause B7, wherein the charge control unit is further configured to: provide an indication to the ground charge sub-control unit indicating which of the plurality of battery packs has a fault.

[0105] 9. The control unit according to any one of clauses B6 to B8, wherein the fault status information includes at least one of overcurrent information, short-circuit information, battery pack health status information, or battery pack temperature information.

[0106] 10. The control unit according to any one of clauses B1 to B9, wherein the charge control unit is further configured to: receive temperature status information from each of the plurality of battery packs; and provide a cooling command to the ground sub-control unit based on the temperature status information.

[0107] Clause set C: 1. A method for charging an aircraft, the method comprising: determining, by one or more processors, a target charge level for each of a plurality of battery packs; receiving, by the one or more processors, charge status information from each of the plurality of battery packs; and commanding, by the one or more processors, one of the plurality of battery packs to disconnect from charging when it is determined that the target charge level of the battery pack has been reached.

[0108] 2. The method according to clause C1, the method further comprising: commanding, by the one or more processors, the ground charge sub-control unit to stop charging when it is determined that the target charge level of each of the plurality of battery packs has been reached.

[0109] 3. The method according to clause C1 or C2, wherein determining the target charge level for each of the plurality of battery packs is based on flight information.

[0110] 4. The method according to clause C3, wherein the flight information includes at least one of the following: distance to the next destination, flight time to the next destination, flight mode to the next destination, expected weather conditions, historical battery consumption information, or charging availability information.

[0111] 5. The method according to clause C4, wherein determining the target charge level of at least one battery pack among the plurality of battery packs is different from the target charge level of another battery pack among the plurality of battery packs.

[0112] 6. The method according to any one of clauses C1 to C5, the method further comprising: receiving, by the one or more processors, fault status information from one of the plurality of battery packs; and when determining that the battery pack has failed, commanding, by the one or more processors, the battery pack to disconnect from charging.

[0113] 7. The method according to clause C6, the method further comprising: when determining that the battery pack has failed, commanding, by the one or more processors, the ground charge sub-control unit to stop charging.

[0114] 8. The method according to clause C7, the method further comprising: providing, by the one or more processors, an indication to the ground charge sub-control unit indicating which battery pack among the plurality of battery packs has failed.

[0115] 9. The method according to any one of clauses C6 to C8, wherein the fault status information includes at least one of overcurrent information, short circuit information, battery pack health status information, or battery pack temperature information.

[0116] 10. The method according to any one of clauses C1 to C9, the method further comprising: receiving, by the one or more processors, temperature status information from each of the plurality of battery packs; and providing, by the one or more processors, a cooling command to the ground sub-control unit based on the temperature status information.

[0117] Clauses D to F Emergency transponder detection

[0118] Clause Set D: An emergency transponder detection system for an aircraft, the emergency transponder detection system comprising: at least one electric propulsion unit (EPU); at least one battery pack configured to supply high-voltage power to the at least one EPU, the battery pack including a battery management system; at least one low-voltage wire connected to the at least one battery pack; wherein the battery management system is configured to: receive aircraft motion information; detect that the motion information indicates a potential collision; detect a current loss in the at least one low-voltage wire; and blow a battery pack fuse to disconnect the supply of the high-voltage power.

[0119] 2. The system according to Clause D1, wherein the battery management system blows the battery pack fuse to disconnect the supply of the high-voltage power when it determines that the potential collision occurs before the current loss in the at least one low-voltage wire.

[0120] 3. The system according to Clause D1 or D2, wherein the battery management system is further configured to: receive the mode of the aircraft from the flight control system of the aircraft; determine a false collision detection based on detecting that the aircraft motion information indicates a potential collision, detecting no current loss in the at least one low-voltage wire, and detecting that the aircraft is in the ground mode at the time of the potential collision; and receive new aircraft motion information based on determining the false collision detection.

[0121] 4. The system according to any one of Clauses D1 to D3, wherein the motion information is aircraft acceleration information received from an accelerometer on the aircraft.

[0122] 5. The system according to Clause D4, wherein the potential collision is detected based on an aircraft acceleration exceeding a threshold.

[0123] 6. The system according to any one of Clauses D1 to D5, the system further comprising at least two battery packs, wherein each battery pack includes the battery management system, a connection to the at least one low-voltage wire, and the battery pack fuse.

[0124] 7. The system according to Clause D6, wherein one of the battery management systems blows all the battery pack fuses of the at least two battery packs when it determines that the potential collision occurs before the current loss in the at least one low-voltage wire.

[0125] 8. The system according to any one of Clauses D1 to D7, wherein the at least one low-voltage wire is routed through the tail of the aircraft.

[0126] 9. The system according to any one of clauses D1 to D5, the system further comprising: at least two battery packs, each battery pack including the battery management system and a connection to the at least one low-voltage wire; and wherein the at least one low-voltage wire for the at least two battery packs is bundled together and routed through the tail of the aircraft.

[0127] 10. The system according to any one of clauses D1 to D9, wherein the battery pack fuse is a pyrotechnic fuse.

[0128] Clause set E: A system for aircraft battery management, the system comprising: a battery management system including one or more processors, wherein the one or more processors are configured to: receive aircraft motion information of the aircraft; detect that the motion information indicates a potential collision; detect a current loss in the at least one low-voltage wire; and blow a battery pack fuse of at least one battery pack configured to supply high-voltage power to disconnect the at least one battery pack from supplying the high-voltage power.

[0129] 2. The system according to clause E1, wherein the battery management system is configured to blow the battery pack fuse to disconnect the at least one battery pack from supplying the high-voltage power when it is determined that the current loss in the at least one low-voltage wire occurs before the potential collision.

[0130] 3. The system according to clause E1 or E2, wherein the battery management system is further configured to: receive the mode of the aircraft from the flight control system of the aircraft; determine a false collision detection based on detecting that the aircraft motion information indicates a potential collision, detecting no current loss in the at least one low-voltage wire, and detecting that the aircraft is in a ground mode at the potential collision; and receive new aircraft motion information based on determining the false collision detection.

[0131] 4. The system according to any one of clauses E1 to E3, wherein the motion information is aircraft acceleration information received from an accelerometer on the aircraft.

[0132] 5. The system according to clause E4, wherein the potential collision is detected based on an aircraft acceleration exceeding a threshold.

[0133] 6. The system according to any one of clauses E1 to E5, the system further comprising at least two battery packs, each battery pack including the battery management system, a connection to the at least one low-voltage wire, and the battery pack fuse.

[0134] 7. The system according to clause E6, wherein, when one of the battery management systems determines that the potential collision occurs before the current loss in at least one low-voltage wire, it fuses all the battery pack fuses of the at least two battery packs.

[0135] 8. The system according to any one of clauses E1 to E7, wherein the at least one low-voltage wire is arranged to pass through the tail of the aircraft.

[0136] 9. The system according to any one of clauses E1 to E5, the system further comprising: at least two battery packs, wherein each battery pack includes the battery management system and a connection to the at least one low-voltage wire; and wherein the at least one low-voltage wire for the at least two battery packs is bundled together and arranged to pass through the tail of the aircraft.

[0137] 10. The system according to any one of clauses E1 to E9, wherein the battery pack fuse is a pyrotechnic fuse.

[0138] Clause set F: A method for aircraft battery management, the method comprising: receiving aircraft motion information by a battery management system; detecting by the battery management system that the motion information indicates a potential collision; detecting by the battery management system a current loss in at least one low-voltage wire; and fusing by the battery management system a battery pack fuse of the at least one battery pack configured to supply high-voltage power to disconnect the supply of high-voltage power from the at least one battery pack.

[0139] 2. The method according to clause F1, the method further comprising fusing the battery pack fuse to disconnect the supply of high-voltage power from the at least one battery pack when it is determined that the potential collision occurs before the current loss in at least one low-voltage wire.

[0140] 3. The method according to clause F1 or F2, the method further comprising: receiving by the battery management system the mode of the aircraft from the flight control system of the aircraft; determining by the battery management system a false collision detection based on detecting that the aircraft motion information indicates a potential collision, detecting no current loss in the at least one low-voltage wire, and detecting that the aircraft is in the ground mode during the potential collision; and receiving by the battery management system new aircraft motion information based on determining the false collision detection.

[0141] 4. The method according to any one of clauses F1 to F3, wherein the motion information is aircraft acceleration information received from an accelerometer on the aircraft.

[0142] 5. The method according to any one of clauses F1 to F4, wherein the potential collision is detected based on an aircraft acceleration exceeding a threshold.

[0143] 6. The method according to any one of clauses F1 to F5, the method further comprising the battery management system communicating with a second battery management system.

[0144] 7. The method according to clause F6, wherein the at least one battery pack includes a first battery pack and a second battery pack configured to supply the high-voltage power, and the battery management system fuses all the battery pack fuses of the first battery pack and the second battery pack before determining a current loss in at least one low-voltage wire when the potential collision occurs.

[0145] 8. The method according to any one of clauses F1 to F7, wherein the battery pack fuse is a pyrotechnic fuse.

[0146] Clause set G Charging infrastructure

[0147] Clause set G: A charging system for an aircraft, the charging system comprising: a plurality of electric propulsion units (EPU); a plurality of battery packs configured to power the plurality of EPUs; a charging port configured to receive high-voltage power to charge the plurality of battery packs; and a common high-voltage charging bus connected to the charging port; wherein the plurality of battery packs are charged through the common high-voltage charging bus; and wherein each of the plurality of battery packs includes a disconnect device to disconnect the battery pack from charging.

[0148] 2. The system according to clause G1, wherein the common high-voltage charging bus is electrically isolated from the high-voltage line powering the plurality of EPUs.

[0149] 3. The system according to clause G1 or G2, the system further comprising: a high-voltage channel at each battery pack, wherein the high-voltage channel connects the battery pack to the common high-voltage charging bus.

[0150] 4. The system according to clause G3, wherein the disconnect device of each battery pack is located on the high-voltage channel.

[0151] 5. The system according to clause G4, wherein the disconnect device includes a contactor.

[0152] 6. The system according to clause G5, wherein the disconnect device includes contactors located on the positive and negative sides of the high-voltage channel.

[0153] 7. The system according to any one of clauses G1 to G6, wherein the plurality of EPUs includes all of the EPUs on one wing of the aircraft.

[0154] 8. The system according to any one of clauses G1 to G7, wherein the charging port is located on the fuselage of the aircraft.

[0155] 9. The system according to any one of clauses G1 to G8, wherein the charging port is further configured to receive communication from a ground charging subsystem configured to supply the high voltage power to charge the plurality of battery packs.

[0156] 10. The system according to clause G9, wherein the charging port is further configured to receive communication from a charging control unit inside the aircraft.

[0157] Clause set H HV architecture:

[0158] 1. A power distribution system for an aircraft, the power distribution system comprising: a plurality of electric propulsion units (EPUs); a first paired battery pack unit including a first battery electrically connected to a second battery via a first high voltage bus, wherein the first battery is configured to supply power to a first group of the plurality of EPUs, and wherein the second battery is configured to supply power to a second group of the plurality of EPUs; and a second paired battery pack unit including a third battery electrically connected to a fourth battery via a second high voltage bus, wherein the third battery is configured to supply power to a third group of the plurality of EPUs, and wherein the fourth battery is configured to supply power to a fourth group of the plurality of EPUs; wherein the first high voltage bus and the second high voltage bus are electrically isolated from each other.

[0159] 2. The system according to clause H1, wherein: the first battery is configured to act as a backup battery for supplying power to the second group of EPUs via the first high voltage bus; the second battery is configured to act as a backup battery for supplying power to the first group of EPUs via the first high voltage bus; the third battery is configured to act as a backup battery for supplying power to the fourth group of EPUs via the second high voltage bus; the fourth battery is configured to act as a backup battery for supplying power to the third group of EPUs via the second high voltage bus;

[0160] 3. The system according to clause H1 or H2, wherein the first high-voltage bus includes a first pyrotechnic fuse, and when the first pyrotechnic fuse is activated, the power of the first group of EPUs is separated from the power of the second group of EPUs.

[0161] 4. The system according to clause H3, wherein the second high-voltage bus includes a second pyrotechnic fuse, and when the second pyrotechnic fuse is activated, the power of the third group of EPUs is separated from the power of the fourth group of EPUs.

[0162] 5. The system according to any one of clauses H1 to H4, wherein each battery includes a pyrotechnic fuse, and each pyrotechnic fuse is configured to disconnect the battery from the operating part of the corresponding battery pack unit in the paired battery pack unit.

[0163] 6. The system according to any one of clauses H1 to H5, wherein the high-voltage lines of the first paired battery pack unit and the high-voltage lines of the second paired battery pack unit are electrically separated from each other.

[0164] 7. The system according to any one of clauses H1 to H6, wherein each EPU of the plurality of EPUs includes a fuse for the high-voltage power supply, and when the fuse is activated, the associated EPU is disconnected from the operating part of the aircraft.

[0165] 8. The system according to any one of clauses H1 to H7, wherein each battery includes a charging contactor, and when the charging contactor is disconnected, the associated battery is disconnected from the high-voltage charging bus.

[0166] 9. The system according to any one of clauses H1 to H8, wherein each group of the groups of EPUs includes two EPUs.

[0167] 10. The system according to clause H9, wherein each group of the groups of EPUs includes two diagonally symmetric EPUs.

[0168] 11. The system according to clause H10, wherein the first paired battery pack unit is configured to supply power to the front outer EPU, the front inner EPU, the rear outer EPU, and the rear inner EPU.

[0169] 12. The system according to clause H11, wherein the second paired battery pack unit is configured to supply power to the front outer EPU, the front inner EPU, the rear outer EPU, and the rear inner EPU.

[0170] 13. The system according to any one of clauses H1 to H12, the system further comprising: a third paired battery pack unit, the third paired battery pack unit including a fifth battery electrically connected to a sixth battery via a third high-voltage bus, wherein the fifth battery is configured to supply power to a fifth group of the plurality of EPUs, and wherein the sixth battery is configured to supply power to a sixth group of the plurality of EPUs; wherein the third high-voltage bus is electrically separated from the first high-voltage bus and the second high-voltage bus.

[0171] 14. The system according to clause H13, wherein: the fifth battery is configured to act as a backup battery for supplying power to the sixth group of EPUs via the third high-voltage bus; and the sixth battery is configured to act as a backup battery for supplying power to the fifth group of EPUs via the third high-voltage bus.

[0172] 15. A power distribution system for an aircraft, the power distribution system comprising: a plurality of electric propulsion units (EPUs); a first paired battery pack unit, the first paired battery pack unit including a first battery, a second battery, and a third battery, the first battery, the second battery, and the third battery being electrically connected via a first high-voltage bus, wherein the first battery is configured to supply power to a first group of the plurality of EPUs, wherein the second battery is configured to supply power to a second group of the plurality of EPUs, and wherein the third battery is configured to supply power to a third group of the plurality of EPUs; and a second paired battery pack unit, the second paired battery pack unit including a fourth battery, a fifth battery, and a sixth battery, the fourth battery, the fifth battery, and the sixth battery being electrically connected via a second high-voltage bus, wherein the fourth battery is configured to supply power to a fourth group of the plurality of EPUs, wherein the fifth battery is configured to supply power to a fifth group of the plurality of EPUs, and wherein the sixth battery is configured to supply power to a sixth group of the plurality of EPUs; wherein the first high-voltage bus and the second high-voltage bus are electrically separated from each other.

[0173] 16. A power distribution system for an aircraft, the power distribution system comprising: a plurality of electric propulsion units (EPUs); a first paired battery pack unit, the first paired battery pack unit including a first battery, a second battery, a third battery, and a fourth battery, the first battery, the second battery, the third battery, and the fourth battery being electrically connected via a first high-voltage bus, wherein the first battery is configured to supply power to a first group of the plurality of EPUs, wherein the second battery is configured to supply power to a second group of the plurality of EPUs, wherein the third battery is configured to supply power to a third group of the plurality of EPUs, and wherein the fourth battery is configured to supply power to a fourth group of the plurality of EPUs; and a second paired battery pack unit, the second paired battery pack unit including a fifth battery, a sixth battery, a seventh battery, and an eighth battery, the fifth battery, the sixth battery, the seventh battery, and the eighth battery being electrically connected via a second high-voltage bus, wherein the fifth battery is configured to supply power to a fifth group of the plurality of EPUs, wherein the sixth battery is configured to supply power to a sixth group of the plurality of EPUs, wherein the seventh battery is configured to supply power to a seventh group of the plurality of EPUs, and wherein the eighth battery is configured to supply power to an eighth group of the plurality of EPUs; wherein the first high-voltage bus and the second high-voltage bus are electrically separated from each other.

[0174] Clause set I General battery management system:

[0175] 1. A system for battery management on a vehicle, the system comprising:

[0176] a first battery pack; a second battery pack; a third battery pack; a first battery management system; a first paired battery pack unit, the first paired battery pack unit including the first battery pack electrically connected to the second battery pack via a high-voltage bus, wherein the first battery pack is configured to supply power to a first electric engine, wherein the second battery pack is configured to supply power to a second electric engine; wherein the first paired battery pack unit is electrically separated from the third battery pack configured to supply power to a third electric engine; wherein the first battery pack is configured to act as a backup battery pack for supplying power to the second electric engine via the high-voltage bus; wherein the second battery pack is configured to act as a backup battery for supplying power to the first electric engine via the high-voltage bus; wherein the first battery management system detects an electrical problem and blows a fuse.

[0177] 2. The system according to clause I1, wherein the electrical problem is an overcurrent condition or a short-circuit condition.

[0178] 3. The system according to Clause I2, wherein the electrical problem is associated with the first electric engine, and fusing the fuse disconnects the first electric engine from the power supply.

[0179] 4. The system according to Clause I2, wherein the electrical problem is associated with the high-voltage bus, and the fusing of the fuse separates the power supply to the first electric engine from the power supply to the second electric engine.

[0180] 5. The system according to Clause I2, wherein the electrical problem is associated with the first battery pack circuit, and the fusing of the fuse cuts off the power supply from the first battery pack to the first electric engine and the second electric engine.

[0181] 6. The system according to any one of Clauses I1 to I5, wherein the first battery management system monitors the charge level of the first battery pack and transmits information about the charge level to the charge control unit.

[0182] 7. The system according to Clause I6, wherein the first battery management system monitors the temperature of the first battery pack and transmits information about the temperature of the first battery pack to the charge control unit.

[0183] 8. The system according to Clause I7, wherein the system further includes a contactor, and the first battery management system disconnects the contactor upon receiving a signal from the charge control unit, disconnecting the first battery pack from the charging circuit.

[0184] 9. The system according to any one of Clauses I1 to I8, wherein the fuse is a pyrotechnic fuse.

[0185] 10. The system according to any one of Clauses I1 to I9, wherein the first electric engine and the second electric engine are electric propulsion units for an electric aircraft.

[0186] Clause set J HV charging mode :

[0187] A control system for charging an aircraft, the control system comprising: a battery pack; an input device, wherein the input device is configured to enable a user to select between different charging modes; two main contactors that connect the battery pack to an electric propulsion unit (EPU) load and an auxiliary load; an EPU load contactor that connects the battery pack to the EPU load; and a controller configured to receive the selected charging mode and control the contactors; wherein the controller is configured to keep the two main contactors open when receiving a user selection to charge in a first mode, thereby disconnecting the EPU load and the auxiliary load; wherein the controller is configured to close the two main contactors and keep the EPU load contactor open when receiving a user selection to charge in a second mode, thereby connecting the auxiliary load and disconnecting the EPU load; and wherein the controller is configured to close the two main contactors and the EPU load contactor when receiving a user selection to charge in a third mode, thereby connecting the auxiliary load and the EPU load.

Claims

1. A charging system for an aircraft, the charging system comprising: A plurality of electric propulsion units (EPU); A plurality of battery packs configured to supply power to the plurality of EPUs; A charging control unit configured to: Determine a target charge level for each of the plurality of battery packs; Receive charge status information from each of the plurality of battery packs; and Command one of the plurality of battery packs to disconnect charging when it is determined that the target charge level of the battery pack has been reached.

2. The system according to claim 1, wherein The charging control unit is further configured to: Command the ground charging subsystem to stop charging when it is determined that the target charge level of each of the plurality of battery packs has been reached.

3. The system according to claim 1 or 2, wherein The charging control unit determines the target charge level for each of the plurality of battery packs based on flight information.

4. The system according to claim 3, wherein, The flight information includes at least one of the following: distance to the next destination, flight time to the next destination, flight mode to the next destination, expected weather conditions, historical battery consumption information, or charging availability information.

5. The system according to claim 4, wherein, Determine that the target charge level of at least one of the plurality of battery packs is different from the target charge level of another of the plurality of battery packs.

6. The system according to any one of claims 1 to 5, wherein, The charging control unit is further configured to: Receive fault status information from each of the plurality of battery packs; and Command one of the plurality of battery packs to disconnect charging when it is determined that the battery pack has failed.

7. The system according to claim 6, wherein, The charging control unit is further configured to: Command the ground charging subsystem to stop charging when it is determined that the battery pack has failed.

8. The system according to claim 7, wherein, The charging control unit is further configured to: Provide an indication to the ground charging subsystem indicating which of the plurality of battery packs has failed.

9. The system according to any one of claims 6 to 8, wherein, The fault status information includes at least one of overcurrent information, short circuit information, battery pack health status information, or battery pack temperature information.

10. The system according to any one of claims 1 to 9, wherein, The charging control unit is further configured to: Receive temperature status information from each of the plurality of battery packs; and Provide a cooling command to the ground subsystem based on the temperature status information.

11. A control unit for charging an aircraft, the control unit comprising: A charging control unit configured to: Determine a target charge level for each of a plurality of battery packs; Receive charge status information from each of the plurality of battery packs; and Command one of the plurality of battery packs to disconnect charging when it is determined that the target charge level of the battery pack has been reached.

12. The control unit according to claim 11, wherein, The charging control unit is further configured to: Command the ground charging control unit to stop charging the plurality of battery packs when it is determined that the target charge level of each of the plurality of battery packs has been reached.

13. The control unit according to claim 11 or 12, wherein, The charging control unit determines the target charge level for each of the plurality of battery packs based on flight information.

14. The control unit according to claim 13, wherein, The flight information includes at least one of the following: distance to the next destination, flight time to the next destination, flight mode to the next destination, expected weather conditions, historical battery consumption information, or charging availability information.

15. The control unit according to claim 14, wherein, Determine that the target charge level of at least one battery pack among the multiple battery packs is different from the target charge level of another battery pack among the multiple battery packs.

16. The control unit according to any one of claims 11 to 15, wherein, The charge control unit is further configured to: Receive fault status information from each of the multiple battery packs; When it is determined that a battery pack has failed, command one of the multiple battery packs to disconnect from charging.

17. The control unit according to claim 16, wherein, The charge control unit is further configured to: when it is determined that a battery pack has failed, command the ground charge sub-control unit to stop charging the battery pack.

18. The control unit according to claim 17, wherein, The charge control unit is further configured to: provide an indication to the ground charge sub-control unit indicating which of the multiple battery packs has failed.

19. The control unit according to any one of claims 16 to 18, wherein, The fault status information includes at least one of overcurrent information, short-circuit information, battery pack health status information, or battery pack temperature information.

20. The control unit according to any one of claims 11 to 19, wherein The charge control unit is further configured to: receive temperature status information from each of the multiple battery packs; and provide a cooling command to the ground sub-control unit based on the temperature status information.

21. A method for charging an aircraft, the method comprising: Determine, by one or more processors, the target charge level of each of the multiple battery packs; Receive, by the one or more processors, charge status information from each of the multiple battery packs; And When it is determined that the target charge level of the battery pack has been reached, command, by the one or more processors, one of the multiple battery packs to disconnect from charging.

22. The method according to claim 21, the method further comprising: When it is determined that the target charge level of each of the multiple battery packs has been reached, command, by the one or more processors, the ground charge sub-control unit to stop charging the multiple battery packs.

23. The method according to claim 21 or 22, wherein Determine that the target charge level of each of the multiple battery packs is based on flight information.

24. The method according to claim 23, wherein, The flight information includes at least one of the following: distance to the next destination, flight time to the next destination, flight mode to the next destination, expected weather conditions, historical battery consumption information, or charging availability information.

25. The method according to claim 24, wherein, Determine that the target charge level of at least one battery pack among the multiple battery packs is different from the target charge level of another battery pack among the multiple battery packs.

26. The method according to any one of claims 21 to 25, the method further comprising: Receive, by the one or more processors, fault status information from one of the multiple battery packs; And when it is determined that the battery pack has failed, command, by the one or more processors, the battery pack to disconnect from charging.

27. The method according to claim 26, the method further comprising: When it is determined that the battery pack has failed, command, by the one or more processors, the ground charge sub-control unit to stop charging.

28. The method according to claim 27, the method further comprising: Provide, by the one or more processors, an indication to the ground charge sub-control unit indicating which of the multiple battery packs has failed.

29. The method according to any one of claims 26 to 28, wherein, The fault status information includes at least one of overcurrent information, short - circuit information, battery pack health status information, or battery pack temperature information.

30. The method according to any one of claims 21 to 29, the method further comprising: Receiving temperature status information from each of the plurality of battery packs by the one or more processors; And providing a cooling command to the ground sub - control unit by the one or more processors based on the temperature status information.

31. An emergency transponder detection system for an aircraft, the emergency transponder detection system comprising: At least one electric propulsion unit (EPU); At least one battery pack configured to supply high - voltage power to the at least one EPU, the battery pack including a battery management system; At least one low - voltage wire connected to the at least one battery pack; Wherein the battery management system is configured to: receive aircraft motion information; detect that the motion information indicates a potential collision; detect a current loss in the at least one low - voltage wire; and blow a battery pack fuse to disconnect the supply of the high - voltage power.

32. The system according to claim 31, wherein, The battery management system blows the battery pack fuse to disconnect the supply of the high - voltage power when determining that the potential collision occurs before the current loss in the at least one low - voltage wire.

33. The system according to claim 31 or 32, wherein The battery management system is further configured to: Receive the mode of the aircraft from the flight control system of the aircraft; Determine a false collision detection based on detecting that the aircraft motion information indicates a potential collision, detecting no current loss in the at least one low - voltage wire, and detecting that the aircraft is in the ground mode at the time of the potential collision; And After determining the false collision detection, receive new aircraft motion information.

34. The system according to any one of claims 31 to 33, wherein The motion information is aircraft acceleration information received from an accelerometer on the aircraft.

35. The system according to claim 34, wherein, Detect the potential collision based on an aircraft acceleration exceeding a threshold.

36. The system according to any one of claims 31 to 35, the system further comprising at least two battery packs, wherein each battery pack includes the battery management system, a connection to the at least one low - voltage wire, and the battery pack fuse.

37. The system according to claim 36, wherein, One of the battery management systems blows all of the battery pack fuses of the at least two battery packs when determining that the potential collision occurs before the current loss in the at least one low - voltage wire.

38. The system according to any one of claims 31 to 37, wherein, The at least one low - voltage wire is routed through the tail of the aircraft.

39. The system according to any one of claims 31 to 35, the system further comprising: At least two battery packs, wherein each battery pack includes the battery management system and the connection to the at least one low - voltage wire; And Wherein the at least one low - voltage wire for the at least two battery packs is bundled together and routed through the tail of the aircraft.

40. The system according to any one of claims 31 to 39, wherein, The battery pack fuse is a pyrotechnic fuse.

41. An aircraft battery management system, the aircraft battery management system comprising: A battery management system including one or more processors, wherein the one or more processors are configured to: Receive aircraft motion information of the aircraft; It is detected that the motion information indicates a potential collision; Detect a current loss in the at least one low-voltage wire; And Fuse a battery pack fuse of at least one battery pack configured to provide high-voltage power to disconnect the supply of the high-voltage power from the at least one battery pack.

42. The system according to claim 41, wherein, The battery management system fuses the battery pack fuse to disconnect the supply of the high-voltage power when it is determined that the potential collision occurs before the current loss in the at least one low-voltage wire.

43. The system according to claim 41 or 42, wherein, The battery management system is further configured to: Receive the mode of the aircraft from the flight control system of the aircraft; Based on detecting that the aircraft motion information indicates a potential collision, detecting no current loss in the at least one low-voltage wire, and detecting that the aircraft is in the ground mode at the potential collision, determine a false collision detection; And After determining the false collision detection, receive new aircraft motion information.

44. The system according to any one of claims 41 to 43, wherein, The motion information is aircraft acceleration information received from an accelerometer on the aircraft.

45. The system according to claim 44, wherein Detect the potential collision based on the aircraft acceleration exceeding a threshold.

46. The system according to any one of claims 41 to 45, the system further comprising at least two battery packs, wherein each battery pack includes the battery management system, a connection to the at least one low-voltage wire, and the battery pack fuse.

47. The system according to claim 46, wherein, One of the battery management systems fuses all the battery pack fuses of the at least two battery packs when it is determined that the potential collision occurs before the current loss in the at least one low-voltage wire.

48. The system according to any one of claims 41 to 47, wherein The at least one low-voltage wire is routed through the tail of the aircraft.

49. The system according to any one of claims 41 to 45, the system further comprising: At least two battery packs, wherein each battery pack includes the battery management system and a connection to the at least one low-voltage wire; and wherein the at least one low-voltage wire for the at least two battery packs is bundled together and routed through the tail of the aircraft.

50. The system according to any one of claims 41 to 49, wherein, The battery pack fuse is a pyrotechnic fuse.

51. A method for aircraft battery management, the method comprising: Receiving, by a battery management system, aircraft motion information of an aircraft; Detecting, by the battery management system, that the motion information indicates a potential collision; Detecting, by the battery management system, a current loss in at least one low-voltage wire; And Fusing, by the battery management system, a battery pack fuse of at least one battery pack configured to provide high-voltage power to disconnect the supply of the high-voltage power from the at least one battery pack.

52. The method according to claim 51, the method further comprising fusing the battery pack fuse to disconnect the supply of the high-voltage power when it is determined that the potential collision occurs before the current loss in the at least one low-voltage wire.

53. The method according to claim 51 or 52, the method further comprising: Receiving, by the battery management system, the mode of the aircraft from the flight control system of the aircraft; The battery management system determines a false collision detection based on detecting that the aircraft motion information indicates a potential collision, detecting no current loss in the at least one low-voltage wire, and detecting that the aircraft is in a ground mode at the potential collision; and after determining the false collision detection, the battery management system receives new aircraft motion information.

54. The method according to any one of claims 51 to 53, wherein, The motion information is aircraft acceleration information received from an accelerometer on the aircraft.

55. The method according to any one of claims 51 to 54, wherein The potential collision is detected based on an aircraft acceleration exceeding a threshold.

56. The method according to any one of claims 51 to 55, the method further comprising the battery management system communicating with a second battery management system.

57. The method according to claim 56, wherein, The at least one battery pack includes a first battery pack and a second battery pack configured to supply the high-voltage power, wherein the battery management system blows all the battery pack fuses of the first battery pack and the second battery pack when determining that the current loss in the at least one low-voltage wire occurs before the potential collision.

58. The method according to any one of claims 51 to 57, wherein The battery pack fuse is a pyrotechnic fuse.

59. A charging system for an aircraft, the charging system comprising: a plurality of electric propulsion units (EPU); a plurality of battery packs configured to power the plurality of EPUs; a charging port configured to receive high-voltage power to charge the plurality of battery packs; and a common high-voltage charging bus connected to the charging port; wherein the plurality of battery packs are charged through the common high-voltage charging bus; and wherein each of the plurality of battery packs includes a disconnect device to disconnect the battery pack from charging.

60. The system according to claim 59, wherein, The common high-voltage charging bus is electrically isolated from the high-voltage line powering the plurality of EPUs.

61. The system according to claim 59 or 60, the system further comprising: A high-voltage channel at each battery pack, wherein the high-voltage channel connects the battery pack to the common high-voltage charging bus.

62. The system according to claim 61, wherein, The disconnect device of each battery pack is located on the high-voltage channel.

63. The system according to claim 62, wherein, The disconnect device includes a contactor.

64. The system according to claim 63, wherein, The disconnect device includes contactors located on the positive and negative sides of the high-voltage channel.

65. The system according to any one of claims 59 to 64, wherein, The plurality of EPUs includes all the EPUs on one wing of the aircraft.

66. The system according to any one of claims 59 to 66, wherein, The charging port is located on the fuselage of the aircraft.

67. The system according to any one of claims 59 to 66, wherein, The charging port is further configured to receive communication from a ground charging subsystem configured to supply the high-voltage power to charge the plurality of battery packs.

68. The system according to claim 67, wherein, The charging port is further configured to receive communication from a charging control unit inside the aircraft.

69. A power distribution system for an aircraft, the power distribution system comprising: a plurality of electric propulsion units (EPU); a first paired battery pack unit including a first battery electrically connected to a second battery via a first high-voltage bus, wherein the first battery is configured to supply power to a first group of the plurality of EPUs, and wherein the second battery is configured to supply power to a second group of the plurality of EPUs; and A second paired battery pack unit, the second paired battery pack unit including a third battery electrically connected to a fourth battery via a second high-voltage bus, wherein the third battery is configured to supply power to a third group of EPUs among the plurality of EPUs, and wherein the fourth battery is configured to supply power to a fourth group of EPUs among the plurality of EPUs; wherein the first high-voltage bus and the second high-voltage bus are electrically separated from each other.

70. The system according to claim 69, wherein: The first battery is configured to act as a backup battery for supplying power to the second group of EPUs via the first high-voltage bus; The second battery is configured to act as a backup battery for supplying power to the first group of EPUs via the first high-voltage bus; The third battery is configured to act as a backup battery for supplying power to the fourth group of EPUs via the second high-voltage bus; The fourth battery is configured to act as a backup battery for supplying power to the third group of EPUs via the second high-voltage bus.

71. The system according to claim 69 or 70, wherein, The first high-voltage bus includes a first pyrotechnic fuse, and upon activation of the first pyrotechnic fuse, the power of the first group of EPUs is separated from the power of the second group of EPUs.

72. The system according to claim 71, wherein, The second high-voltage bus includes a second pyrotechnic fuse, and upon activation of the second pyrotechnic fuse, the power of the third group of EPUs is separated from the power of the fourth group of EPUs.

73. The system according to any one of claims 69 to 72, wherein, Each battery includes a pyrotechnic fuse, and each pyrotechnic fuse is configured to disconnect the battery from the operating portion of the corresponding one of the paired battery pack units.

74. The system according to any one of claims 69 to 73, wherein, The high-voltage lines of the first paired battery pack unit and the high-voltage lines of the second paired battery pack unit are electrically separated from each other.

75. The system according to any one of claims 69 to 74, wherein, Each EPU of the plurality of EPUs includes a fuse for the high-voltage power supply, wherein upon activation of the fuse, the associated EPU is disconnected from the operating portion of the aircraft.

76. The system according to any one of claims 69 to 75, wherein, Each battery includes a charging contactor, and when the charging contactor is disconnected, the associated battery is disconnected from the high-voltage charging bus.

77. The system according to any one of claims 69 to 76, wherein, Each group of the groups of EPUs includes two EPUs.

78. The system according to claim 77, wherein, Each group of the groups of EPUs includes two diagonally symmetric EPUs.

79. The system according to claim 78, wherein, The first paired battery pack unit is configured to supply power to the front outer EPUs, the front inner EPUs, the rear outer EPUs, and the rear inner EPUs.

80. The system according to claim 79, wherein, The second paired battery pack unit is configured to supply power to the front outer EPUs, the front inner EPUs, the rear outer EPUs, and the rear inner EPUs.

81. The system according to any one of claims 69 to 80, the system further comprising: A third paired battery pack unit, the third paired battery pack unit including a fifth battery electrically connected to a sixth battery via a third high-voltage bus, wherein the fifth battery is configured to supply power to a fifth group of EPUs among the plurality of EPUs, and wherein the sixth battery is configured to supply power to a sixth group of EPUs among the plurality of EPUs; wherein the third high-voltage bus is electrically separated from the first high-voltage bus and the second high-voltage bus.

82. The system according to claim 81, wherein: The fifth battery is configured to act as a backup battery for powering the sixth set of EPUs via the third high-voltage bus; and the sixth battery is configured to act as a backup battery for powering the fifth set of EPUs via the third high-voltage bus.

83. A power distribution system for an aircraft, the power distribution system comprising: A plurality of electric propulsion units (EPUs); A first paired battery pack unit including a first battery, a second battery, and a third battery, the first battery, the second battery, and the third battery being electrically connected via a first high-voltage bus, wherein the first battery is configured to supply power to a first set of the plurality of EPUs, wherein the second battery is configured to supply power to a second set of the plurality of EPUs, and wherein the third battery is configured to supply power to a third set of the plurality of EPUs; And A second paired battery pack unit including a fourth battery, a fifth battery, and a sixth battery, the fourth battery, the fifth battery, and the sixth battery being electrically connected via a second high-voltage bus, wherein the fourth battery is configured to supply power to a fourth set of the plurality of EPUs, wherein the fifth battery is configured to supply power to a fifth set of the plurality of EPUs, and wherein the sixth battery is configured to supply power to a sixth set of the plurality of EPUs; Wherein the first high-voltage bus and the second high-voltage bus are electrically isolated from each other.

84. A power distribution system for an aircraft, the power distribution system comprising: A plurality of electric propulsion units (EPUs); A first paired battery pack unit including a first battery, a second battery, a third battery, and a fourth battery, the first battery, the second battery, the third battery, and the fourth battery being electrically connected via a first high-voltage bus, wherein the first battery is configured to supply power to a first set of the plurality of EPUs, wherein the second battery is configured to supply power to a second set of the plurality of EPUs, wherein the third battery is configured to supply power to a third set of the plurality of EPUs, and wherein the fourth battery is configured to supply power to a fourth set of the plurality of EPUs; and A second paired battery pack unit including a fifth battery, a sixth battery, a seventh battery, and an eighth battery, the fifth battery, the sixth battery, the seventh battery, and the eighth battery being electrically connected via a second high-voltage bus, wherein the fifth battery is configured to supply power to a fifth set of the plurality of EPUs, wherein the sixth battery is configured to supply power to a sixth set of the plurality of EPUs, wherein the seventh battery is configured to supply power to a seventh set of the plurality of EPUs, and wherein the eighth battery is configured to supply power to an eighth set of the plurality of EPUs; Wherein the first high-voltage bus and the second high-voltage bus are electrically isolated from each other.

85. A system for battery management on a vehicle, the system comprising: A first battery pack; A second battery pack; A third battery pack; A first battery management system; A first paired battery pack unit, the first paired battery pack unit comprising the first battery pack electrically connected to the second battery pack via a high-voltage bus, wherein the first battery pack is configured to supply power to a first electric engine, and wherein the second battery pack is configured to supply power to a second electric engine; Wherein the first paired battery pack unit is electrically separated from the third battery pack configured to supply power to a third electric engine; Wherein the first battery pack is configured to act as a backup battery pack for supplying power to the second electric engine via the high-voltage bus; Wherein the second battery pack is configured to act as a backup battery for supplying power to the first electric engine via the high-voltage bus; Wherein the first battery management system detects an electrical problem and blows a fuse.

86. The system according to claim 85, wherein, The electrical problem is an overcurrent condition or a short-circuit condition.

87. The system according to claim 86, wherein, The electrical problem is associated with the first electric engine, and the blowing of the fuse disconnects the first electric engine from the power source.

88. The system according to claim 86, wherein, The electrical problem is associated with the high-voltage bus, and the blowing of the fuse separates the power supply to the first electric engine from the power supply to the second electric engine.

89. The system according to claim 86, wherein, The electrical problem is associated with the first battery pack circuit, and the blowing of the fuse cuts off the power supply from the first battery pack to the first electric engine and the second electric engine.

90. The system according to any one of claims 85 to 89, wherein, The first battery management system monitors the charge level of the first battery pack and transmits information about the charge level to a charge control unit.

91. The system according to claim 90, wherein, The first battery management system monitors the temperature of the first battery pack and transmits information about the temperature of the first battery pack to the charge control unit.

92. The system according to claim 91, wherein, The system further includes a contactor, and the first battery management system disconnects the contactor upon receiving a signal from the charge control unit to disconnect the first battery pack from the charging circuit.

93. The system according to any one of claims 85 to 92, wherein, The fuse is a pyrotechnic fuse.

94. The system according to any one of claims 85 to 93, wherein, The first electric engine and the second electric engine are electric propulsion units for an electric aircraft.

95. A control system for charging an aircraft, the control system comprising: A battery pack; An input device configured to enable a user to select between different charging modes; Two main contactors that connect the battery pack to an electric propulsion unit (EPU) load and an auxiliary load; An EPU load contactor that connects the battery pack to the EPU load; And A controller configured to receive the selected charging mode and control the contactors; Wherein the controller is configured to keep the two main contactors open upon receiving a user selection to charge in a first mode, thereby disconnecting the connection to the EPU load and the auxiliary load; Wherein, the controller is configured to close the two main contactors and keep the EPU load contactor open when receiving a user selection to charge in the second mode, thereby connecting the auxiliary load and disconnecting the EPU load; And wherein the controller is configured to close the two main contactors and the EPU load contactor when receiving a user selection to charge in the third mode, thereby connecting the auxiliary load and the EPU load.

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