High Voltage Battery Architecture

Abstract

A power distribution system for an aircraft with multiple electric propeller units (EPUs) 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 multiple EPUs. The system includes a second paired battery pack 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 set of EPUs and a fourth set of EPUs, respectively, of the multiple EPUs. The first high-voltage bus and the second high-voltage bus are electrically separate from each other.

Description

[Technical Field]

[0001] CROSS-REFERENCE TO RELATED APPLICATIONS This disclosure claims priority to and benefit of U.S. Provisional Application No. 63 / 383,660, filed November 14, 2022, entitled "Systems and Methods for Improved Battery Assemblies for eVTOL Aircraft" (Attorney Docket No. 16163.6005-00000), the contents of which are incorporated herein in their entirety for all purposes.

[0002] This disclosure relates generally to the field of powered air vehicles. More specifically, and without limitation, this disclosure relates to innovations in tiltrotor aircraft using electric propulsion systems. Certain aspects of this disclosure generally relate to the configuration and control of high-voltage power supply systems for aircraft. [Background technology]

[0003] Electric aircraft include battery packs that power various flight components, including the electric propulsion units (EPUs) that enable flight. These battery packs are critical to enabling the EPUs to provide lift and thrust support for the aircraft. Therefore, there is a need to provide redundancy in the aircraft's power systems to avoid single points of failure. There is also a need to prevent a fault or failure condition from cascading to and damaging other critical aircraft components. The disclosed high-voltage power system solves these and other problems by connecting battery packs together in battery pack units, where each battery pack within a unit serves as a backup for the other battery packs. Furthermore, each battery pack unit is electrically separate from the other battery pack units.

[0004] Additionally, to ensure that the battery packs can power the EPU for the duration of the flight, the battery packs must be fully charged before takeoff. Therefore, the battery packs must be charged efficiently and effectively. The disclosed high-voltage power system solves these and other problems by controlling the amount of charge to the battery packs based on upcoming flight information, historical battery pack information, and monitored battery pack conditions. The disclosed high-voltage power system also solves these and other problems by providing a single charging point for multiple battery packs.

[0005] Finally, in the event of a crash, there is a need to enable first responders to 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 loop connected to the battery pack. When a first responder detects that the low-voltage cut loop has been cut, a fuse to the battery pack is blown and the high-voltage power system is de-energized. The cut loop may be routed in the tail of the aircraft to provide isolation from the high-voltage lines and increase the safety of first responders. Summary of the Invention

[0006] The present disclosure generally relates to an electric power system for an aircraft. One aspect of the present disclosure provides a power distribution system for an aircraft including a plurality of electric propeller units (EPUs). The system includes a first paired battery pack unit including a first battery electrically connected to a second battery via a first high-voltage bus, where the first battery is configured to provide power to a first set of EPUs of the plurality of EPUs, and the second battery is configured to provide power to a second set of EPUs of the plurality of EPUs. The system includes a second paired battery pack unit including a third battery electrically connected to a fourth battery via a second high-voltage bus, where the third battery is configured to provide power to the third set of EPUs of the plurality of EPUs, and the fourth battery is configured to provide power to the fourth set of EPUs of the plurality of EPUs. Furthermore, the first high-voltage bus and the second high-voltage bus are electrically separate from each other. [Brief explanation of the drawings]

[0007] [Figure 1a] 1 illustrates an exemplary eVTOL aircraft consistent with embodiments of the present disclosure. [Figure 1b] 1 illustrates another exemplary eVTOL aircraft consistent with embodiments of the present disclosure. [Figure 1c] 1 illustrates an electric engine 110 having two partial motors consistent with an embodiment of the present disclosure. [Figure 1d] 1 illustrates a diagram of a high-voltage power distribution system for an eVTOL aircraft, consistent with an embodiment of the present disclosure. [Figure 2a] 1 illustrates a circuit diagram of a high voltage junction box (HVJB) consistent with an embodiment of the present disclosure. [Figure 2b] 1 illustrates a diagram of a high voltage junction box (HVJB) consistent with an embodiment of the present disclosure. [Figure 3] 1 illustrates a view of a charge port assembly (CPA) consistent with an embodiment of the present disclosure. [Figure 4] 1 illustrates a flowchart for detecting emergency responders, consistent with an embodiment of the present disclosure. [Figure 5a] 10 illustrates a plan view for routing cut loop wiring through the tail of an eVTOL aircraft, consistent with an embodiment of the present disclosure. [Figure 5b] FIG. 10 illustrates a profile diagram for routing cut loop wiring through the tail of an eVTOL aircraft, consistent with an embodiment of the present disclosure. DETAILED DESCRIPTION OF THE INVENTION

[0008] The present disclosure relates primarily to components of electric vertical take-off and landing (eVTOL) aircraft used in non-traditional aircraft. For example, an eVTOL aircraft of the present disclosure may be intended for frequent (e.g., 50 or more flights per workday) short-duration flights (e.g., less than 100 miles per flight) to and from densely populated areas. The aircraft may be intended to carry four to six passengers or commuters who expect a low-noise and low-vibration experience. Therefore, it is desirable that these components be configured and designed for frequent use without wear, generate little heat or vibration, and include mechanisms to effectively control and manage the heat or vibration generated by the components. Also, several of these aircraft may be intended to operate near each other in congested urban areas. Therefore, it may be desirable that these components be configured and designed to generate low levels of noise both inside and outside the aircraft and include various safety and backup mechanisms. For example, for safety reasons, it may be desirable that the aircraft be propelled by a distributed propulsion system to avoid the risk of a single point of failure and be capable of conventional takeoff and landing on a runway. Furthermore, it may be desirable for an aircraft to be able to safely take off and land vertically from a relatively confined space (e.g., a vertiport, parking lot, or driveway) compared to a conventional airport runway, and to be able to transport approximately four to six passengers or commuters and their associated luggage. These usage requirements may impose design constraints on the size, weight, and operational efficiency (e.g., drag, energy use) of the aircraft, which may affect the design and configuration of the aircraft's components.

[0009] The disclosed embodiments provide new and improved aircraft component configurations not observed in conventional aircraft and / or identified design criteria that differ from conventional aircraft components. Such alternative configurations and design criteria are combined to address shortcomings and challenges of conventional components, resulting in the embodiments disclosed herein for eVTOL aircraft components of various configurations and designs.

[0010] In some embodiments, the disclosed eVTOL aircraft is designed for both vertical and conventional takeoff and landing, enabling vertical flight, forward flight, and transitions via a distributed electric propulsion system. Thrust may be generated by supplying high-voltage power to the distributed electric propulsion system's electric engines, each of which may convert the high-voltage power into mechanical shaft power for rotating a propeller. The embodiments disclosed herein may contribute to optimizing the energy density of the electric propulsion system. Embodiments may include an electric engine connected to an onboard power source, which may include a device capable of storing energy, such as a battery or capacitor, or one or more systems for harnessing or generating electricity, such as a fuel-powered generator or a solar panel array. Some disclosed embodiments provide weight and space savings for aircraft components, thereby improving aircraft efficiency and performance. Focusing on passenger transport safety, the disclosed embodiments implement new and improved safety protocols and system redundancy in the event of a malfunction, minimizing single points of failure in the aircraft propulsion system. Some disclosed embodiments provide new and improved approaches to meeting aviation and transportation laws and regulations.

[0011] FIG. 1A illustrates an exemplary eVTOL aircraft consistent with embodiments of the present disclosure. As shown in FIG. 1A, in some embodiments, the distributed electric propulsion system of the eVTOL aircraft 100 may include twelve electric engines 110, which may be mounted on forward and aft booms of the aircraft's 100 wings. The forward electric engine 110 may be tiltable during flight between a horizontally oriented position (e.g., to generate forward thrust) and a vertically oriented position (e.g., to generate vertical lift). The forward electric engine 110 may be of a clockwise or counterclockwise type relative to the direction of propeller rotation. The aft electric engine 110 may be fixed in a vertically oriented position (e.g., to generate vertical lift) and may also be of a clockwise or counterclockwise type relative to the direction of propeller rotation.

[0012] The aircraft 100 may have various combinations of forward and aft electric engines 110. For example, in some embodiments, the aircraft 100 may have six forward electric engines 110 and six aft electric engines 110. In some other embodiments, the aircraft 100 may include four forward electric engines 110 and four aft electric engines 110, or any other combination of forward engines 110 and aft engines 110. In some other embodiments, the number of forward electric engines and aft electric engines is not equal.

[0013] In some embodiments, for vertical takeoff and landing (VTOL) missions, the forward electric engine 110 and the aft electric engine 110 may provide vertical thrust during takeoff and landing. During flight phases when the aircraft 100 is in forward flight mode, the forward electric engine 110 may provide horizontal thrust, while the propeller of the aft electric engine 110 may be stowed in a fixed position to minimize drag. The aft electric engine 110 may be actively stowed with position monitoring.

[0014] In some embodiments, for conventional takeoff and landing (CTOL) missions, the forward electric engine 110 may provide horizontal thrust for fixed-wing takeoff, cruise, and landing. In some embodiments, the aft electric engine 110 need not be used to generate thrust during CTOL missions, and the aft propeller may be stowed in place.

[0015] Transition from vertical flight to forward flight and vice versa may be achieved via a tilt propeller subsystem that can redistribute thrust from primarily vertical during vertical flight mode to primarily horizontal during forward flight mode. A variable pitch mechanism may vary the collective angle of the propeller hub assembly blades of the forward electric engine for operation during hover, transition, and cruise phases.

[0016] The tilting propeller system may include a linear or rotary actuator for changing the orientation of the propulsion system during operation. In some embodiments, the pitch of the propulsion system may be changed as a function of the orientation of the propulsion system. In some embodiments, the rotary actuator may include a motor, an inverter, and a gearbox. In some embodiments, the gearbox may include various types of gears that interface to provide a gear reduction that can orient the propulsion system. In some embodiments, the tilting propeller system may include a redundant configuration in which multiple motors, inverters, and gearboxes are present and interface using gears. In some embodiments, a configuration utilizing multiple motors, gearboxes, and inverters may allow a failed portion of the redundant configuration to be driven by a motor, inverter, and gearbox of another portion of the configuration. In some embodiments, the gearbox configuration may also allow the tilting propeller system to maintain the orientation of the propulsion system with or without the aid of additional power provided by the system.

[0017] In some embodiments, the electric engine 110 may be housed in or connected to the boom of the aircraft 100 and may include a motor, an inverter, and a gearbox. In some embodiments, the motor, inverter, and gearbox may be interfaced such that they share a central axis. In some embodiments, torque due to the motor may be sent to the gearbox, away from the propeller of the propulsion system. In some embodiments, the gearbox may provide a gear reduction and then send torque back to the propeller via the main shaft through bearings located inside the motor. In some embodiments, the inverter may be mounted to the rear of the gearbox so that the main shaft does not move past the inverter when outputting torque to the propeller.

[0018] As shown in FIG. 1A, the aircraft 100 may be configured with a distributed electric propulsion system that enables vertical flight, forward flight, and transition. The forward six electric engines 110 (numbered 1 through 6 from left to right) tilt controllable pitch propellers to achieve vertical takeoff and landing, transition flight, and full fixed-wing flight. The aft six electric engines 110 (numbered 7 through 12 from left to right) are equipped with fixed pitch propellers that operate during vertical takeoff and landing and transition, and are stowed within a minimum drag position for conventional flight. Flight control is an integrated fly-by-wire system featuring envelope protection and structural load limiting capabilities. The aircraft 100 is equipped with advanced cockpit avionics, a flight management system, and sensors necessary to support intended operation and system function.

[0019] In some embodiments, the electric propulsion system (EPS) described herein may generate thrust by supplying high-voltage (HV) electrical power to electric engines 110, which in turn convert the HV electrical power into mechanical shaft power used to rotate propellers. As mentioned above, the aircraft 100 described herein may have multiple electric engines 110 boom-mounted forward and aft of the wings. The amount of thrust generated by each electric engine 110 may be controlled by torque commands from a flight control system (FCS) via a digital communication interface to each electric engine 110. Embodiments may include forward electric engines 110, which may be capable of changing their orientation, or cant. Additional embodiments include forward engines, which may be of a clockwise (CW) or counterclockwise (CCW) type. The forward electric engine propulsion subsystem may consist of a multi-blade adjustable pitch propeller as well as a variable pitch subsystem.

[0020] In some embodiments, aircraft 100 includes a high-voltage power supply (HVPS) system for providing high-voltage (HV) electrical power. The HVPS system is a source of electrical power on aircraft 100 and is configured to convert electrical power into mechanical rotating shaft power and distribute stored electrical energy to other systems on aircraft 100, including an electric propulsion system (EPS) for generating thrust. As shown in FIG. 1a, aircraft 100's HVPS system may include six battery packs 120 (numbered 1 through 6 from left to right) installed in battery bays in the wings of 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 power one or more electric engines 110. Although six battery packs 120 are shown, aircraft 100 may have any number of battery packs 120.

[0021] In some embodiments, a single battery pack 120 may be electrically connected to and power multiple electric engines 110. For example, in some embodiments, battery pack 120 may power electric engines 110 on either side of a longitudinal axis. In some embodiments, battery pack 120 may power electric engines 110 on either side of a horizontal axis. In some embodiments, as shown in FIG. 1a, battery pack 120 may power two diagonally opposed electric engines 110. For example, battery pack 1 may power electric engines 1 and 12. Battery pack 2 may power electric engines 5 and 8. Battery pack 3 may power electric engines 3 and 10. Battery pack 4 may power electric engines 4 and 9. Battery pack 5 may power electric engines 2 and 11. Battery pack 6 may power electric engines 6 and 7. Thus, upon loss of battery pack 120, impacts to roll or pitch moments may be reduced because the loss of lift is balanced. In some embodiments, battery pack 120 may power different arrangements of electric engines 110 to reduce roll, pitch, or yaw moments that may be caused by the loss of battery pack 120. For example, in some embodiments, battery pack 120 may be connected to electric engines 110 in any manner that balances lift and / or thrust across the longitudinal and horizontal axes of the aircraft.

[0022] Additionally, the HVPS system includes a cross-link 130 having at least one fuse that enables pairing of two or more battery packs 120. Through the cross-link, power for the electric engine 110 can be shared between the paired battery packs 120. Thus, multiple battery packs 120 can simultaneously power multiple electric engines 110. This arrangement provides redundancy and avoids a single point of failure because each paired battery 120 can serve as a backup for the other battery(ies). In the event of a battery pack 120 failure, one or more connected battery packs 120 can continue to power the electric engine 110 connected to the failed battery pack.

[0023] 1a, 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, which provides power to electric engines 1 and 12, may be cross-linked to battery pack 4, which provides power to electric engines 4 and 9. Battery pack 2, which provides power to electric engines 5 and 8, may be cross-linked to battery pack 5, which provides power to electric engines 2 and 11. Battery pack 3, which provides power to electric engines 3 and 10, may be cross-linked to battery pack 6, which provides power to electric engines 6 and 7.

[0024] FIG. 1b illustrates another exemplary eVTOL aircraft consistent with embodiments of the present disclosure. In some embodiments, electric engine 110 may include multiple motor stages, each 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 can 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 a first partial motor on electric engines 1, 6, 7, and 12. Battery pack 6 may power a second partial motor on electric engines 1, 6, 7, and 12. In some embodiments, different configurations may be used. For example, two battery packs may provide power to the partial motors on electric engines 1, 4, 9, and 12.

[0025] 1c illustrates an electric engine 110 having two partial motors 191a and 191b consistent with embodiments of the present disclosure. The 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 or may operate simultaneously to drive the blades at higher power. The partial motors 191a and 191b are driven by their own motor controllers 192a and 192b, respectively. In some embodiments, the power to the partial motors may be electrically separate such that each electric engine 110 has an electrically separate backup.

[0026] The above configurations are provided as examples, but different numbers and configurations of battery packs 120, electric engines 110, battery pack to electric engine connections, and battery pack cross-link combinations may be used. In some embodiments, each battery pack 120 may power an individual electric engine 110. For example, an aircraft may have 4, 6, 8, 10, 12, 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 separate 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 supply, but the battery packs 120 are still electrically separate.

[0027] In some embodiments, each battery pack 120 may power multiple electric engines 110. As described above, battery packs 120 may power a set of electric engines 110 that are symmetrical across one or more axes of symmetry. In some embodiments, battery packs 120 may power electric engines 110 that are symmetrical across a longitudinal axis, a 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.

[0028] In some embodiments, battery pack 120 may power more than two electric engines 110. In some embodiments, battery pack 120 may power two or more sets of diagonally symmetric electric engines. For example, in some embodiments, 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 battery pack 120 may include an inner diagonally symmetric pair of electric engines 110 and an outer diagonally symmetric pair of electric engines 110.

[0029] In some embodiments, battery pack 120 may power four or more electric engines 110 in a configuration that is symmetrical across a longitudinal axis of symmetry. For example, battery pack 120 may power electric engines 1, 6, 7, and 12. In some embodiments, in each of the above configurations, battery pack 120 may provide 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 supply, but battery pack 120 remains electrically separate.

[0030] In some embodiments, some or all of the battery packs 120 are interconnected. As described above, cross-links 130 may allow each battery pack 120 to serve as backup power for another battery pack. For example, in some embodiments, battery pack 1 may directly power a first number of electric engines, and a second battery pack 120 may directly power a second number of electric engines. The first and second battery packs 120 may be cross-linked together to form a battery pack unit. Thus, each battery pack in the unit may function as a backup for the other battery packs. In the event of a failure of a battery pack in the unit, the failed battery pack may be disconnected, and the electric engine 110 may be powered by one or more surviving battery packs in the unit. The battery packs in a battery pack unit may be electrically separate from the other battery pack units.

[0031] As described above, in some embodiments, a battery pack unit may include two battery packs 120, each battery pack 120 powering several electric engines 110. As described above, in some embodiments, each battery pack 120 may power two diagonally symmetric electric engines 110. Thus, each battery pack unit may power a total of four electric engines 110, each electric engine having a battery pack backup. In some embodiments, each battery pack 120 in a battery pack unit may power four electric engines 110, comprising two diagonally symmetric sets of electric engines 110. Thus, each battery pack unit may power a total of eight electric engines 110, each electric engine having a battery pack backup.

[0032] In some embodiments, a battery pack unit may include three battery packs 120, each battery pack powering several electric engines 110. For example, in some embodiments, each battery pack 120 may power two diagonally symmetric electric engines 110. Thus, each battery pack unit may power a total of six electric engines 110, with each electric engine 110 having two battery pack backups. In some embodiments, each battery pack 120 in a battery pack unit may power four electric engines 110, comprising two diagonally symmetric sets of electric engines 110. Thus, each battery pack unit may power a total of 12 electric engines 110, with each electric engine having two battery pack backups.

[0033] In some embodiments, a battery pack unit may include four battery packs 120, each battery pack powering several electric engines. For example, in some embodiments, each battery pack may power two diagonally symmetric electric engines 110. Thus, each battery pack unit may power a total of eight electric engines 110, each electric engine having three battery pack backups. In other embodiments, each battery pack 120 in a battery pack unit may power four electric engines 110, comprising two diagonally symmetric sets of electric engines 110. Thus, each battery pack unit may power 16 electric engines 110, each electric engine having three battery pack backups.

[0034] In some embodiments, electric engine 110 includes a single motor powered by one or more battery packs 120. In some embodiments, each electric engine 110 may include two or more partial motors, and battery pack 120 may power the partial motors. In some embodiments, the power supply configuration of electric engine 110 described above may include battery packs powering the partial motors. For example, in some embodiments, each electric engine 110 may include two partial motors, and a battery pack unit may power the partial motors of electric engines 2, 4, 6, 7, 9, and 11. A second battery pack unit may power the partial motors of electric engines 2, 3, 6, 7, 10, and 11. A third battery pack unit may power the partial motors of electric engines 1, 3, 5, 8, 10, and 12. A fourth battery pack unit may power the partial motors of electric engines 1, 4, 5, 8, 9, and 12. Thus, each electric engine 110 receives backup power through the other partial motors. As described above, each battery pack unit may include one or more battery packs, for example, a battery pack unit may include one, two, three, or four battery packs.

[0035] In some embodiments, each electric engine may include two partial motors, and a battery pack unit may power the partial motors of electric engines 2, 3, 4, 5, 8, 9, 10, and 11. A second battery pack unit may power the partial motors of electric engines 1, 2, 5, 6, 7, 8, 11, and 12. A third battery pack unit may power the partial motors of electric engines 1, 3, 4, 6, 7, 9, 10, and 12. Thus, each electric engine 110 receives backup power through the other partial motors. As described above, each battery pack unit may include one or more battery packs. For example, a battery pack unit may include one, two, three, or four battery packs.

[0036] In some embodiments, each electric engine may include two partial motors, and a battery pack unit may power the partial motors of electric engines 3, 4, 9, and 10. A second battery pack unit may power the partial motors of electric engines 3, 4, 9, and 10. A third battery pack unit may power the partial motors of electric engines 2, 5, 8, and 11. A fourth battery pack unit may power the partial motors of electric engines 2, 5, 8, and 11. A fifth battery pack unit may power the partial motors of electric engines 1, 6, 7, and 12. A sixth battery pack unit may power the partial motors of electric engines 1, 6, 7, and 12. Thus, each electric engine 110 receives backup power through the other partial motors. As described above, each battery pack unit may include one or more battery packs. For example, a battery pack unit may include one, two, three, or four battery packs. Different configurations of battery packs 120, electric engines 110, battery pack-to-electric engine connections, and battery pack cross-link combinations can be selected to optimally balance aircraft power needs, system redundancy, and fault tolerance.

[0037] FIG. 1d illustrates a diagram of a high-voltage power supply system for an eVTOL aircraft consistent with embodiments of the present disclosure. As shown in FIG. 1b, the eVTOL aircraft may include a battery assembly comprising electrically separate battery pack units (e.g., 160, 162, and 164). Each battery pack unit may include a battery pack 120 that is cross-linked together, as described above. In some embodiments, the battery pack units may include battery packs 120 that ensure aircraft controllability is maintained upon loss of a battery pack unit. Thus, upon loss of a battery pack unit, the aircraft may still be controllable. As described above, in some embodiments, the battery pack units may include battery packs 120 that power electric engines 110 on either side of one or more axes of symmetry. Thus, upon loss of a battery pack unit, impacts to roll, pitch, or yaw moments may be reduced because the loss of lift and / or thrust is balanced. In some embodiments, the loss of power or reduction in power caused by a failure of a battery pack unit has a substantially symmetric effect with respect to the roll, pitch, and / or yaw of the aircraft (e.g., asymmetry of <±5%, <±10%, <±15%, <±20%, or <±25%). In some embodiments, the battery pack unit may include battery packs 120 that reduce the amount of high-voltage wiring between the battery packs. In some embodiments, the battery pack unit may include battery packs 120 to minimize power requirements.

[0038] In some embodiments, as shown in FIG. 1d, the HVPS system may include three electrically separate battery pack units. For example, in some embodiments, battery pack unit 160 may include battery packs 1 and 4 powering electric engines 1, 4, 9, and 12. Battery pack unit 162 may include battery packs 2 and 5 powering electric engines 2, 5, 8, and 11. Battery pack unit 164 may include battery packs 3 and 6 powering 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. In the event of a failure of one battery pack 120 in a battery pack unit, the other paired battery pack 120 continues to power the four electric engines.

[0039] In some embodiments, each battery pack unit 160, 162, 164 may include a high-voltage bus for cross-linking the battery packs 120 within the battery pack unit. In some embodiments, the cross-link 130 connects two high-voltage channels, each supplying one or more electric engines 110. For example, in some embodiments, the 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., powering two electric engines). The cross-link may further include a bus that connects the negative-voltage channels after they are combined (e.g., after powering two electric engines).

[0040] In some embodiments, each cross link 130 may include at least one fuse for disconnecting the cross link in the event of a cross link failure. For example, fuses 131, 132, and 134 may be located on the cross link connections of the positive high voltage channels in battery pack units 160, 162, and 164. In some embodiments, the fuses may be pyrotechnical fuses. As described in further detail below, the battery management system of the connected battery pack 120 may determine a fault in the cross link, such as a short circuit or overcurrent condition, and blow the associated pyrotechnical fuse. Thus, the cross link may be disconnected, avoiding further damage to HVPS system components (e.g., electric engine, battery, EPUS). Furthermore, the electric engine 110 will still receive power from the paired battery pack 120 in the battery pack unit. For example, in the event of a cross-link failure, pyrotechnical fuse 131 may be blown, but electric engines 1 and 12 still receive power from battery pack 1, and electric engines 4 and 9 still receive power from battery pack 4.

[0041] In some embodiments, additional pyrotechnical fuses may be present on the cross-link connections of the negative high-voltage channels. For example, pyrotechnical fuses 170, 172, and 174 may be located on the cross-link connections in battery pack units 160, 162, and 164, respectively. This configuration may provide additional redundancy to the system. If the fuse on the positive cross-link connection fails, the fuse on the negative cross-link connection may act as a backup, and vice versa. For example, in some embodiments, if the fuse on the positive cross-link connection does not blow after being commanded, the connected battery management system may instruct the negative cross-link fuse to blow. Furthermore, in some embodiments, each positive cross-link may have two fuses controlled by the two associated battery packs, and each negative cross-link may have two fuses controlled by the two associated battery packs.

[0042] In some embodiments, the HVPS system may include load disconnection devices for disconnecting portions of the HVPS circuitry upon a fault (e.g., a short circuit or overcurrent condition) of a downstream electric engine, downstream EPU, or other downstream distribution circuitry. In some embodiments, the load disconnection devices may be located directly upstream of the electric engines. For example, in some embodiments, load disconnection devices 109, 111, 112, and 113 may be located on the high-voltage channels powering engines 1, 12, 4, and 9, respectively. Load disconnection devices 114, 115, 116, and 117 may be located on the high-voltage channels powering engines 2, 11, 5, and 8, respectively. Load disconnection devices 118, 119, 121, and 122 may be located on the high-voltage channels powering engines 3, 10, 6, and 7, respectively.

[0043] In some embodiments, the load disconnection device is a pyrotechnical fuse. In the event of a failure of a downstream component, the pyrotechnical fuse may receive a signal (e.g., from the battery management system of the connected battery) and blow the fuse. Thus, the downstream component may be disconnected, and further damage to other equipment (e.g., the electric engine, the battery, the EPUS) may be avoided. Furthermore, the remaining electric engine 110 in the battery pack unit will still receive power from the connected battery pack 120. For example, in the event of a failure of a device or wiring downstream of the pyrotechnical fuse 109, the pyrotechnical fuse 109 may blow, but the electric engines 12, 4, and 9 will still receive power from the battery packs 1 and 4. Furthermore, in some embodiments, the load disconnection device may include a contactor, and the battery management system may instruct the contactor to disconnect the circuit. In some embodiments, both the contactor and the fuse may be used to provide additional redundancy, and the pyrotechnical fuse may act as a backup for the contactor.

[0044] 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 charge disconnect device. In some embodiments, the charge 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 provide disconnection for battery packs 1, 4, 5, 2, 3, and 6, respectively. Similarly, in some embodiments, additional charge 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 provide disconnection for battery packs 1, 4, 5, 2, 3, and 6, respectively.

[0045] In some embodiments, the charge disconnect device is a contactor, such as K4 positive and K4 negative in FIG. 2A . The charge contactor may function as a redundant measure to disconnect battery pack 120 from charging. As described in further detail below, battery pack 120 may report a charging problem to a charge control unit (CCU). For example, battery pack 120 may report a short circuit or overcurrent condition within battery pack 120 or within the high-voltage charging channel. In some embodiments, if the CCU fails to stop charging, battery pack 120 may instruct the charge contactor to disconnect the charging channel. In some embodiments, battery pack 120 may automatically instruct the charge contactor to disconnect the charging channel without waiting for the CCU to fail. In some embodiments, after instructing the CCU to stop charging and / or disconnecting the battery pack 120 that detected a charging problem, battery pack 120 and / or the CCU may instruct other battery packs 120 to disconnect from the charging channel. Disconnecting battery pack 120 when a charging problem is detected can avoid damage to HVPS components.

[0046] FIG. 2a illustrates 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 loads 210 to provide high-voltage power. Specifically, a DC / DC converter and a storage element BT1 (e.g., parallel and series connected battery cells) in a battery management system (BMS) may be used to provide the high-voltage power. The DC / DC converter and the storage element BT1 are connected to each of the HV loads through a combination of pre-charge resistor(s) (e.g., resistor R1) or current-sensing resistor(s) (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) for protection against various fault conditions (e.g., overcurrent, short circuit, etc.). In some embodiments, the fuses F1-F7 may be one or more of the fuses detailed above with respect to FIG. 1b. For example, in some embodiments, fuses F2 EE1, F3 EE2, and F4 Xlink may correspond to fuses 109, 111, and 131 detailed in FIG. 1d.

[0047] Fuse F1 may be a pack fuse for disconnecting a failed battery pack 120 from the rest of the HVPS system. In some embodiments, F1 may be a pyrotechnical fuse. Upon failure of battery pack 120, pyrotechnical fuse F1 may receive a signal (e.g., from an associated battery management system) and blow fuse F1. Thus, further damage to other equipment (e.g., electric engine, EPU, connected battery packs) may be avoided. Furthermore, electric engine 110 will still receive power from the paired battery pack 120 in the battery pack unit. For example, upon battery pack failure, pyrotechnical fuse F1 of battery pack 1 may be blown, but electric engines 1, 12, 4, and 9 may still receive power from battery pack 4.

[0048] The circuit arrangement within the high-voltage junction box (HVJB) 222 provides charging flexibility by allowing auxiliary loads and / or the electric engine and actuators to be energized or de-energized during the charging process. For example, the battery pack 120 can be charged while the remaining HVPS circuits remain disconnected. Charging contactors K4 positive and K4 negative can be closed to allow the battery pack 120 to charge. Meanwhile, main contactors K1 and K2 and pre-charge contactors (and / or relays) K3 and K5 can be opened to prevent the remaining HVPS circuits from being energized. Furthermore, the battery pack 120 can be charged while the auxiliary loads are connected but the electric engine and actuators remain disconnected. Charging contactors K4 positive and K4 negative can be closed to allow the battery pack 120 to charge. Meanwhile, main contactors K1 and K2 can be closed after pre-charge contactor (and / or relay) K3 has finished pre-charging the auxiliary loads, and K4EE can remain open. Furthermore, 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. Meanwhile, the main contactors K1 and K2 can be closed and K4EE can be closed after the pre-charge contactors (and / or relays) K3 and K5 have finished pre-charging the connected loads.

[0049] In some embodiments, the input device may allow a person to select a charging mode for the aircraft. For example, a person may request charging in one of the three different modes outlined above through the input device. In some embodiments, the input device may be a physical switch, button, and / or lever. In some embodiments, the input device may be a user interface element provided on a display screen or control panel. In some embodiments, the input device may be a processor that can receive manual selection and / or voice command requesting a mode switch. The input device may include any means that allows a person to select a desired charging mode. In some embodiments, the input information is transmitted to the BMS 270, which may control the contactors according to the requested charging mode.

[0050] FIG. 2b illustrates a diagram of a high-voltage junction box 222 (HVJB) consistent with embodiments of the present disclosure. In some embodiments, each battery pack 120 houses an HV distribution unit 211, a battery management system (BMS 270), and a pyro-fuse redundant trigger board (PRT 280) housed within the HVJB 122. Each unit may be a hardware device such as a computer, processor, or microprocessor. The BMS 270 may be configured to monitor voltage, temperature, current, and isolation resistance. The BMS 270 may control battery pack contactors and pyrotechnical fuses to protect against fault conditions. As described in further detail below, the BMS 270 may communicate with various systems within and outside the HVJB 222. The BMS 270 may include a battery management unit (BMU 271) that may receive voltage, current, resistance, and temperature sensing signals from the cell stack assembly 224 and / or the HV distribution unit 211. The BMS 270 may further include a cell management unit (CMU) 272 for monitoring the voltage of each set of seven parallel cells connected in series within the 14S-7P cell block (i.e., a 1S-7P cell group). The CMU may also be used to monitor the temperature of the 14S-7P cell block. The CMU 272 obtains measurements for all cell groups within the battery pack 120 and communicates the measurements to the BMU 271.

[0051] The BMU 271 may monitor the output current for each connected load. The BMU 271 may be internally powered by the battery cell stack assembly 224 and may continuously monitor the status of the batteries, even when the batteries are not installed in the aircraft 100. By monitoring battery pack 120, cell block, and cell group parameters, the BMU can protect against conditions that adversely affect safety or performance, such as overvoltage, undervoltage, overheating, underheating, loss of electrical insulation, short circuits, and overcurrent. The diagnostic capabilities of the BMU 271 enable fault detection and isolation through built-in tests (BITs). Additionally, the BMU 271 performs calculations of the battery pack 120's state of charge (SOC), state of health (SOH), fault conditions (e.g., short circuit or overcurrent), state of power (SOP), state of energy (SOE), and state of temperature (SOT). The BMU 271 also controls and monitors bus precharge, provides fuse and contactor commands, and communicates with various systems within and outside the HVJB 222.

[0052] The HV distribution unit 211 within the HVJB 222 may house an HV contactor 212 and a combination of active and passive fuses (e.g., pyrotechnical fuses 213 and 214) to protect against overcurrent and short-circuit conditions. In some embodiments, the contactor 212 may correspond to one or more of the switching devices K1-K7 (e.g., HV contactors) detailed in FIG. 2a. Similarly, the pyrotechnical fuses 213 and 214 may correspond to one or more fuses F1-F8 detailed in FIG. 2a. The HV distribution unit 211 may further include (or receive information from) current sensors (e.g., resistors R3-R6, Hall effect sensors, shunt current sensors, or other sensor(s)).

[0053] In some embodiments, a pyrofuse redundant trigger board (PRT 280) may be located within HVJB 222. In other embodiments, BMS 270 may communicate with PRT 280 located outside HVJB 222. BMS 270 may detect a fault event and send a command signal to PRT 280 for a corresponding pyrofuse driver to blow a fuse. For example, in some embodiments, HV distribution unit 211 may receive a sensor signal from a current sensor (e.g., resistors R3-R6) and provide information to BMU 271 regarding the condition (e.g., voltage, current, or temperature) of a connected load at a point in the HVPS system. Based on the received information, BMU 271 may determine a fault condition (e.g., because a value is outside a predetermined range) and send a command to PRT 280 to blow the associated pyrotechnical fuse. Thus, the fault condition can be disconnected from the rest of the HVPS circuitry, protecting the remaining devices and wiring. In some embodiments, the BMU 271 may monitor the sensors directly instead of receiving the information through the HV distribution unit 211 .

[0054] In some embodiments, battery packs 120 may communicate with each other, for example, through BMS 270. Battery packs 120 may use information about the status of one or more paired battery packs 120 within a battery pack unit to help determine if an overcurrent condition has occurred. For example, battery pack 120 may determine an expected operating range (e.g., voltage, current, etc.) based on the battery pack status and the communication status of battery packs 120 within the battery pack unit. In some embodiments, HVJB 122 may further provide a redundant active trigger board configured to enable a pyrofuse driver to activate one or more pyrotechnical fuses when BMS 270 fails to enable the pyrofuse driver. See U.S. Pat. No. 11,710,957, incorporated by reference.

[0055] The control MCU (CCU 263) in the charge port assembly 262 interfaces with the external battery charger and may communicate with the BMUs 271 on the six installed battery packs 120. This unit may be a hardware device such as a computer, processor, or microprocessor. In some embodiments, the CCU 263 may be a single PCBA with one microcontroller managing the overall power delivery to each battery pack 120 during charging. As shown in FIG. 2, the CCU 263 may perform the handshake between the ground charging subsystem 274 and the BMUs 271 and command the BMUs 271 to open and close contactors 212, such as contactors K6-K7 detailed in FIG. 2a. The CCU 263 may perform active detection and protection functions for overvoltage protection. The BMUs 271 in each battery pack 120 retain complete control and may continuously monitor their battery packs 120 during charging operations.

[0056] FIG. 3 illustrates a diagram of a charge port assembly (CPA 262) consistent with an embodiment of the present disclosure. Charge port assembly 262 includes a charge port 330 that provides a communication connection through powerline communication 335 and HV power transmission through an HV power channel 336. In some embodiments, charge port 330 may be a JI772 Type 1 charge port that includes various pins and connection points to enable connection to a ground service system (GSS 300) (e.g., via a plug). Charge port 330 may include one or more proximity pins to detect a high-voltage connection between GSS 300 and charge port 330. Upon detection of a connection with GSS 300, charge port 330 may engage a latch that prevents high-voltage power 336 from being disconnected under a charged load. Following completion of charging, charge port 330 may automatically unlatch the connection or allow manual unlatching.

[0057] The charge port assembly 262 may include a charge control unit (CCU 263) that communicates with the charge port 330, for example, through communication line 333. The CCU 263 may also provide latch control 331, lighting changes 334, and monitor and respond to the temperature 332 of various components. The CCU 263 may monitor the temperature of the inlet side of the charge port 330. If the temperature is too high, the CCU 263 may instruct the ground servicing system 300 to stop charging. The CCU 263 receives status updates from the battery packs 120 and provides instructions to the battery packs 120 to control their charge levels by opening and closing battery pack charging contactors (e.g., K6-K7 in FIG. 2A ). As described in more detail above with reference to FIG. 2b , in some embodiments, the CCU 263 may communicate with the battery management system (BMS 270) of each battery pack, for example, through a battery management unit (BMU 271). The BMS 270 may send battery pack information to the CCU 263, including information regarding the battery pack connection status (e.g., whether the battery pack is connected to the HVPS system), state of charge (SOC), state of health (SOH), fault status (e.g., short circuit or overcurrent), state of power (SOP), state of energy (SOE), and state of temperature (SOT).

[0058] The CCU 263 may provide commands to the BMS 270 to open or close the battery pack charging contactors. In some embodiments, each battery pack 120 may have a separate low-voltage CAN communication line connecting the battery pack 120 to the CCU 263. In some embodiments, a CAN communication line may be shared among one or more battery packs 120 in a battery pack unit. For example, HV battery packs 1 and 4 may communicate with the CCU 263 through CAN 351. HV battery packs 2 and 5 may communicate with the CCU 263 through CAN 352. HV battery packs 3 and 6 may communicate with the CCU 263 through CAN 353. As described in further detail below, the CCU 263 may make various powering and cooling requests of the GSS 300 (e.g., via the charging port 330) based on information received from the battery packs 120.

[0059] Charging control unit 263 may determine battery pack charging contactor commands based on various criteria. In some embodiments, CCU 263 may determine the required battery pack charge level based on flight information. For example, in some embodiments, CCU 263 may receive flight information from GSS 300, for example, through communication lines 335 and 333. GSS 300 may receive flight information through a wired or wireless connection to a computer, laptop, iPad, mobile device, or any other device capable of providing flight information. In some embodiments, charging port assembly 262 may provide a direct wired or wireless connection to a computer, laptop, iPad, or mobile device to receive flight information directly. In some embodiments, CCU 263 may receive flight information from aircraft flight control system 230.

[0060] The flight information may 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 may include the expected type of flight. For example, the flight mission information may include the duration or distance to be covered by each flight mode. In some embodiments, the flight modes may include wing-assisted flight, thrust and lift-assisted flight, thrust-assisted flight, and lift-assisted flight. In some embodiments, the flight mission information may include the expected EPU power throughout the flight, for example, as units of power or a percentage of maximum EPU power. In some embodiments, flight mission information may be provided for each EPU on the aircraft.

[0061] The flight mission information may include information regarding weather conditions expected throughout the flight. Weather conditions may include temperature, pressure, wind conditions, and precipitation expected throughout the flight. The flight mission information may include, for example, an expected weight of the aircraft based on the number of passengers or the amount of cargo. The weight of the aircraft may be predicted or measured (e.g., if the aircraft is carrying and charging passengers or cargo).

[0062] The flight information may include historical battery information. For example, in some embodiments, the battery information may include historical battery consumption for each battery pack on a particular flight path. The battery information may further include details regarding flight mode, weight, and weather for the charge control unit 263 to determine its relevance to upcoming flight missions.

[0063] Additionally, flight information for multiple subsequent flights may be received and analyzed. In some embodiments, if the aircraft makes multiple trips without the ability to recharge, flight information for all subsequent flights may be collected and analyzed to ensure the aircraft has sufficient charge for each trip. In some embodiments, the aircraft may have time to partially recharge before a subsequent trip. Thus, the flight information may include information about the subsequent trips, including the amount of recharge available between trips. By receiving this information, CCU 263 may ensure that battery packs 120 have sufficient charge to support a sufficient portion of the subsequent trips. CCU 263 may use the flight information to determine the required charge level needed for each battery pack 120.

[0064] The charge control unit (CCU 263) may determine battery pack charge contactor commands based on the current status of each battery pack 120 received from the BMS 270, including the state of energy and / or state of charge of each battery pack 120. The CCU 263 may determine the amount of additional charge needed to meet the required charge level based on the current charge level of each battery pack. Additionally, in some embodiments, the CCU 263 may consider the battery pack configuration when charging the battery packs 120. The CCU 263 may determine to charge each battery pack 120 in a battery pack unit to the same charge level. Thus, the CCU 263 may charge all battery packs 120 in a battery pack unit to the highest charge level required for any battery pack 120 in the unit. As the battery packs 120 charge, the CCU 263 may continue to receive updates regarding the charge level of each battery pack and may keep the battery pack's charge contactors closed to allow charging until the required charge level is reached.

[0065] Additionally, the charging control unit (CCU 263) may determine battery pack charging contactor commands based on fault, health, or temperature conditions received from the BMS 270. In some embodiments, the CCU 263 may open the contactors to the battery pack 120 (disable charging) based on receiving information that the battery pack 120 has failed (e.g., experienced a short circuit or overcurrent condition). Additionally, the CCU 263 may open the contactors to the battery pack 120 (disable charging) based on the battery pack's health condition dropping below a set level or the battery pack's temperature exceeding a set level. The CCU 263 may continue to monitor the fault, health, or temperature conditions from the battery pack 120 and close the contactors (enable charging) when the conditions improve.

[0066] Charging control unit 263 may send cooling commands to GSS 300, for example, through communication lines 333 and 335, based on state temperature information received from battery pack 120. In some embodiments, CCU 263 may send a required battery pack temperature or a required coolant flow rate. Ground charging subsystem 310 may communicate this information to thermal regulation subsystem 320. Thermal regulation subsystem 320 may control one or more condensers and associated coolant control valves to achieve the cooling requirements.

[0067] The charging control unit (CCU 263) may signal the status of the battery pack 120 to the charger throughout the charging process. In some embodiments, the CCU 263 may signal a problem (e.g., a battery pack failure, poor health, or excessive temperature) through illumination line 334. For example, in some embodiments, a light may illuminate or change color to indicate a problem. Alternatively or additionally, the CCU 263 may communicate details of the problem (e.g., the type of problem, the associated battery pack(s), etc.) to the ground servicing system 300 through communication lines 333 and 335. The ground servicing system 300 may provide these details through a display, computer, laptop, iPad, mobile device, or any other device capable of communicating information to the charger.

[0068] The charge control unit (CCU 263) may determine when each battery pack 120 has reached the required charge level and may signal charging completion to the ground servicing system 300. Upon determining that charge is not being received from the GSS 300, the CCU 263 may provide a signal to the charge port, for example, through latch control 331, to enable automatic unlatching of the connection to the GSS 300 or manual unlatching of the connection.

[0069] FIG. 4 illustrates a flowchart for detecting emergency responders consistent with embodiments of the present disclosure. In some embodiments, this process may be performed by each battery management system 270 of a battery pack 120. In step 401, a 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 one associated with an aircraft's flight control system. In step 402, the processor receives a high-voltage interlock loop (HVIL) continuity status (e.g., from the battery management system (BMS) 270) indicating whether a low-voltage emergency cut loop has been cut. For example, the BMS 270 may determine that the cut 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. Furthermore, in some embodiments, the collected information may include a timestamp indicating when it was collected, while in other embodiments, the processor may assign a time based on when the information was received.

[0070] In step 403, the processor may determine whether an emergency responder has performed a cut of the low-voltage emergency cut loop. The processor may make this determination based on acceleration information and HVIL continuity status. If the acceleration information indicates a crash (e.g., exceeds a threshold) at an earlier time than the HVIL continuity status indicates a cut loop, it is determined that an emergency responder has cut the low-voltage emergency cut loop. However, if the HVIL continuity status indicates a cut loop at an earlier time than the acceleration information indicates a crash, no emergency response is detected. Furthermore, if the acceleration information does not indicate a crash or the HVIL continuity status does not indicate a cut, no emergency response is detected. If it is determined in step 404 that an emergency responder has performed a cut, 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 loop. For example, in some embodiments, cut loop 1 may be connected to battery pack 1. The processor may determine an emergency responder cut loop 1, and the processor may instruct battery pack 1 to blow a pyrotechnical fuse of battery pack 1, such as fuse F1 in FIG. 2a. In some embodiments, based on determining that an emergency responder cut any of the loops, the processor may blow pyrotechnical fuses associated with battery pack 120 and any connected battery packs 120. For example, referring to FIG. 1a, based on determining that an emergency responder associated with battery pack 1 cut loop 1, the processor may blow pyrotechnical fuses associated with battery packs 1 and 4. In some embodiments, based on determining that an emergency responder cut any of the loops, the processor may blow pyrotechnical fuses associated with all battery packs 120.

[0071] In step 405, the processor may determine whether the collision detection was false. The processor may determine that the acceleration information indicates a collision, but the HVIL continuity status indicates that there is no cut loop. Additionally, the processor may collect or have available information regarding whether the flight control system is in ground mode. If the processor determines that the aircraft's flight control system is in ground mode, the processor may determine that the collision detection was false. However, if the processor determines that the flight control system is not in ground mode (e.g., flight mode), a false collision is not determined. In some embodiments, "ground mode" may be a mode selected by the pilot through an interface when the pilot is operating the aircraft on the ground.

[0072] If it is determined in step 406 that the crash detection was false, condition 1 is reset to indicate that no crash was detected, and the processor collects acceleration information again. If it is determined in step 407 that the crash detection was not false, condition 1 is not reset, and the processor continues to monitor whether the HVIL continuity status indicates a cut loop in condition 2 of step 403.

[0073] FIG. 5a illustrates a plan view for routing cut loop wiring through the tail of an eVTOL aircraft, consistent with embodiments of the present disclosure. As described above, each cut loop may be connected to a single battery pack 120. Thus, six cut loops may be routed from battery packs 120 located in the wings or elsewhere to the tail of the aircraft. This routing ensures that the cut loops are accessible for cutting at the tail of the aircraft, away from the high-voltage power system that runs between the batteries, electric engines, and other aircraft devices that run toward the front of the aircraft. First responders can cut one or more loops to de-energize a battery pack 120 without risking cuts to live high-voltage wires, thereby increasing safety. In some embodiments, each cut loop may be routed separately. In some embodiments, cut loops may be routed with one or more battery packs 120 (e.g., in a bundle). For example, cut loops associated with connected battery packs may be bundled together, or cut loops associated with the wings of the aircraft may be bundled together. In some embodiments, all of the cut loops of the battery pack 120 may be routed together in a single bundle.

[0074] FIG. 5b illustrates a profile diagram for routing cut loop wiring through the tail of an eVTOL aircraft, consistent with embodiments of the present disclosure. As detailed above, the cut loops are routed through the tail of the aircraft to enhance safety for emergency responders. Additionally, the cut loops may be routed in a manner that allows them to be easily accessible by first responders. For example, in some embodiments, the cut loops may be routed toward the perimeter of the aircraft, making them easy to find and cut. The cut loops may be color-coded and include descriptive tags at set intervals to ensure easy identification by first responders.

[0075] The embodiments may be further described using the following clauses. Provision Sets A-C Charging for Aircraft Clause Set A: A charging system for an aircraft, the charging system comprising: a plurality of electric propeller units (EPUs); a plurality of battery packs configured to power the plurality of EPUs; and a charge control unit, the charge 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 a battery pack of the plurality of battery packs to disconnect from charging upon determining that the target charge level for the battery pack has been reached. 2. The system of clause A1, wherein the charge control unit is further configured to instruct the ground charging subsystem to cease charging upon determining that a target charge level for each of the plurality of battery packs has been reached. 3. The system of clause A1 or A2, wherein the charge control unit determines a target charge level for each of the plurality of battery packs based on flight information. 4. The system described in clause A3, wherein the flight information includes at least one of 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 recharge availability information. 5. The system of clause A4, wherein a target charge level for at least one of the plurality of battery packs is determined to be different from a target charge level for another battery pack of the plurality of battery packs. 6. The system of any one of clauses A1-A5, wherein the charging control unit is further configured to receive fault status information from each of the plurality of battery packs and to instruct a battery pack of the plurality of battery packs to disconnect from charging upon determining a fault associated with the battery pack. 7. The system of clause A6, wherein the charging control unit is further configured to command the ground charging subsystem to cease charging upon determining a fault with the battery pack. 8. The system of clause A7, wherein the charging control unit is further configured to provide an indication to the ground charging subsystem of which battery pack of the plurality of battery packs has failed. 9. The system of any one of clauses A6-A8, wherein the fault status information includes at least one of overcurrent information, short circuit information, battery pack health information, or battery pack temperature information. 10. The system of 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 cooling instructions to the ground subsystem based on the temperature status information.

[0076] Clause Set B: A control unit for charging an aircraft, comprising: a charge 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 instruct a battery pack of the plurality of battery packs to disconnect from charging upon determining that the target charge level for the battery pack has been reached. 2. The control unit described in clause B1, wherein the charging control unit is further configured to instruct the ground charging sub-control unit to cease charging the plurality of battery packs upon determining that a target charge level for each of the plurality of battery packs has been reached. 3. The control unit of clause B1 or B2, wherein the charge control unit determines a target charge level for each of the plurality of battery packs based on flight information. 4. The control unit described in clause B3, wherein the flight information includes at least one of 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 recharge availability information. 5. The control unit of clause B4, wherein a target charge level for at least one of the plurality of battery packs is determined to be different from a target charge level for another battery pack of the plurality of battery packs. 6. The charge control unit receives fault status information from each of the plurality of battery packs; The control unit of any one of clauses B1-B5, further configured to command a battery pack of the plurality of battery packs to disconnect from charging upon determining a fault associated with the battery pack. 7. The control unit described in clause B6, wherein the charging control unit is further configured to instruct the ground charging sub-control unit to cease charging the battery pack upon determining a fault associated with the battery pack. 8. The control unit of clause B7, wherein the charging control unit is further configured to provide an indication to the ground charging sub-control unit of which battery pack of the plurality of battery packs has failed. 9. The control unit of any one of clauses B6-B8, wherein the fault status information includes at least one of overcurrent information, short circuit information, battery pack health information, or battery pack temperature information. 10. A control unit described in any one of clauses B1 to B9, wherein the charging control unit is further configured to receive temperature status information from each of the plurality of battery packs and provide cooling instructions to the ground sub-control unit based on the temperature status information.

[0077] Clause Set C:1. A method for charging an aircraft, 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 instructing, by the one or more processors, a battery pack of the plurality of battery packs to disconnect from charging upon determining that the target charge level for the battery pack has been reached. 2. Method clause C1 further including instructing, by the one or more processors, the ground charging sub-control unit to discontinue charging upon determining that a target charge level for each of the plurality of battery packs has been reached. 3. The method of clause C1 or C2, wherein determining the target charge level for each of the plurality of battery packs is based on flight information. 4. The method of clause C3, wherein the flight information includes at least one of 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 recharge availability information. 5. The method of clause C4, wherein the target charge level for at least one of the plurality of battery packs is determined to be different from the target charge level for another battery pack of the plurality of battery packs. 6. The method of any one of clauses C1-C5, further including: receiving, by one or more processors, fault status information from a battery pack of the plurality of battery packs; and instructing, by one or more processors, the battery pack to disconnect from charging upon determining a fault associated with the battery pack. 7. The method of clause C6, further comprising instructing, by the one or more processors, the ground charging sub-control unit to discontinue charging upon determining a fault with the battery pack. 8. Method clause C7 further including providing, by the one or more processors, to the ground charging sub-control unit an indication of which battery pack of the plurality of battery packs has failed. 9. The method of any one of clauses C6-C8, wherein the fault status information includes at least one of overcurrent information, short circuit information, battery pack health information, or battery pack temperature information. 10. The method of any one of clauses C1 to C9, further comprising: receiving, by one or more processors, temperature status information from each of the plurality of battery packs; and providing, by the one or more processors, cooling instructions to the ground sub-control unit based on the temperature status information.

[0078] Articles D-F Emergency Responder Detection Clause Set D: An emergency responder detection system for an aircraft, comprising: at least one electric propeller 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; and at least one low-voltage wire connected to the at least one battery pack, the battery management system configured to receive aircraft motion information, detect that the motion information indicates a potential collision, detect a loss of current in the at least one low-voltage wire, and blow a battery pack fuse to disconnect the supply of high-voltage power. 2. The system of clause D1, wherein the battery management system blows a battery pack fuse to disconnect the supply of high voltage power upon determining a potential crash that occurred prior to a loss of current in at least one low voltage wire. 3. The system described in clause D1 or D2, wherein the battery management system is further configured to receive an aircraft mode from the aircraft flight control system, determine a false collision detection based on detecting that the aircraft motion information indicates a potential collision, detecting that there is no loss of current in at least one low-voltage wire, and detecting that the aircraft was in ground mode at the time of the potential collision, and receive new aircraft motion information based on determining the false collision detection. 4. The system of any one of clauses D1-D3, wherein the motion information is aircraft acceleration information received from an accelerometer on the aircraft. 5. The system of any one of clauses D4, wherein a potential collision is detected based on aircraft acceleration exceeding a threshold. 6. A system described in any one of clauses D1 to D5, further comprising at least two battery packs, each battery pack comprising a battery management system, a connection to at least one low-voltage wire, and a battery pack fuse. 7. The system described in clause D6, wherein one of the battery management systems blows all battery pack fuses of the at least two battery packs upon determining a potential crash occurred prior to a loss of current in at least one low-voltage wire. 8. The system of any one of clauses D1-D7, wherein at least one low voltage wire is routed through the tail of the aircraft. 9. The system of any one of clauses D1-D5, further comprising at least two battery packs, each battery pack comprising a battery management system and a connection to 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. 10. The system of any one of clauses D1-D9, wherein the battery pack fuse is a pyrotechnical fuse.

[0079] Clause Set E: A system for battery management for an aircraft, comprising: a battery management system including one or more processors, the one or more processors configured to: receive aircraft motion information for the aircraft; detect that the motion information indicates a potential collision; detect a loss of current in 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 supply of high-voltage power by the at least one battery pack. 2. The system described in clause E1, wherein the battery management system is configured to blow a battery pack fuse to disconnect the supply of high voltage power by the at least one battery pack upon determining a potential crash occurring prior to a loss of current in the at least one low voltage wire. 3. The system of clause E1 or E2, wherein the battery management system is further configured to receive an aircraft mode from the aircraft flight control system, determine a false collision detection based on detecting that the aircraft motion information indicates a potential collision, detecting that there is no loss of current in at least one low-voltage wire, and detecting that the aircraft was in ground mode at the time of the potential collision, and receive new aircraft motion information based on determining the false collision detection. 4. The system of any one of clauses E1-E3, wherein the motion information is aircraft acceleration information received from an accelerometer on the aircraft. 5. The system of clause E4, wherein a potential collision is detected based on aircraft acceleration exceeding a threshold. 6. The system of any one of clauses E1-E5, further comprising at least two battery packs, each battery pack comprising a battery management system, a connection to at least one low-voltage wire, and a battery pack fuse. 7. The system of any one of clauses E6, wherein one of the battery management systems blows all battery pack fuses of at least two battery packs upon determining a potential crash occurred prior to a loss of current in at least one low-voltage wire. 8. The system of any one of clauses E1-E7, wherein at least one low voltage wire is routed through the tail of the aircraft. 9. The system of any one of clauses E1-E5, further comprising at least two battery packs, each battery pack comprising a battery management system and a connection to 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. 10. The system of any one of clauses E1-E9, wherein the battery pack fuse is a pyrotechnical fuse.

[0080] Clause Set F: A method for aircraft battery management, comprising: receiving, by a battery management system, aircraft motion information; detecting, by the battery management system, that the motion information indicates a potential collision; detecting, by the battery management system, a loss of current in at least one low voltage wire; and blowing, by the battery management system, a battery pack fuse of at least one battery pack configured to supply high voltage power to disconnect the supply of high voltage power by the at least one battery pack. 2. The method of clause F1, further comprising blowing a battery pack fuse to cut off the supply of high voltage power by at least one battery pack upon determining a potential collision occurred prior to a loss of current in at least one low voltage wire. 3. The method of clause F1 or F2, further including receiving, by the battery management system, an aircraft mode from a flight control system of the aircraft; determining, by the battery management system, a false collision detection based on detecting, by the battery management system, that the aircraft motion information indicates a potential collision, detecting that there is no loss of current in at least one low-voltage wire, and detecting that the aircraft was in ground mode at the time of the potential collision; and receiving, by the battery management system, new aircraft motion information based on determining the false collision detection. 4. The method of any one of clauses F1-F3, wherein the motion information is aircraft acceleration information received from an accelerometer on the aircraft. 5. The method of any one of clauses F1-F4, wherein a potential collision is detected based on aircraft acceleration exceeding a threshold. 6. The method of any one of clauses F1-F5, further comprising a battery management system in communication with a second battery management system. 7. The method of clause F6, wherein the at least one battery pack comprises a first battery pack and a second battery pack configured to supply high-voltage power, and the battery management system blows all battery pack fuses of the first battery pack and the second battery pack upon determining a potential collision occurring prior to loss of current in at least one low-voltage wire. 8. The method of any one of clauses F1-F7, wherein the battery pack fuse is a pyrotechnical fuse.

[0081] Clause Set G Charging Infrastructure Clause Set G: A charging system for an aircraft, comprising: a plurality of electric propeller units (EPUs); a plurality of battery packs configured to power the plurality of EPUs; a charging port configured to accept high voltage power for charging the plurality of battery packs; and a common high voltage charging bus connected to the charging ports, wherein the plurality of battery packs are charged through the common high voltage charging bus, and each of the plurality of battery packs includes a disconnect device for disconnecting the battery pack from charging. 2. The system of clause G1, wherein the common high-voltage charging bus is electrically separate from the high-voltage wiring powering the multiple EPUs. 3. The system of clause G1 or G2, further comprising a high voltage channel for each battery pack, the high voltage channel connecting the battery packs to a common high voltage charging bus. 4. The system described in clause G3, wherein the disconnect device for each battery pack is located on the high voltage channel. 5. The system of clause G4, wherein the disconnecting device comprises a contactor. 6. The system described in clause G5, wherein the disconnecting device includes contactors on both the positive and negative sides of the high voltage channel. 7. The system of any one of clauses G1-G6, wherein the multiple EPUs include all EPUs on one wing of the aircraft. 8. The system of any one of clauses G1-G7, wherein the charging port is located on the fuselage of the aircraft. 9. The system of any one of clauses G1-G8, wherein the charging port is further configured to accept communications from a ground charging subsystem configured to supply high-voltage power to charge a plurality of battery packs. 10. The system of clause G9, wherein the charging port is further configured to accept communications from a charging control unit onboard the aircraft.

[0082] Clause Set H HV Architecture 1. A power distribution system for an aircraft, the power distribution system comprising: a plurality of electric propeller units (EPUs); first paired battery pack units, the first paired battery pack unit comprising a first battery electrically connected to a second battery via a first high-voltage bus, the first battery configured to provide power to the EPUs of a first set of the plurality of EPUs, and the second battery configured to provide power to the EPUs of a second set of the plurality of EPUs; and second paired battery pack units, the second paired battery pack unit comprising a third battery electrically connected to a fourth battery via a second high-voltage bus, the third battery configured to provide power to the EPUs of the third set of the plurality of EPUs, and the fourth battery configured to provide power to the EPUs of a fourth set of the plurality of EPUs, wherein the first high-voltage bus and the second high-voltage bus are electrically separate from each other. 2. The system described in clause H1, wherein the first battery is configured to function as a backup battery for powering a second set of EPUs via a first high-voltage bus, the second battery is configured to function as a backup battery for powering the first set of EPUs via the first high-voltage bus, the third battery is configured to function as a backup battery for powering a fourth set of EPUs via the second high-voltage bus, and the fourth battery is configured to function as a backup battery for powering a third set of EPUs via the second high-voltage bus. 3. A system described in clause H1 or H2, wherein the first high voltage bus includes a first pyrotechnical fuse, and power to the first set of EPUs is isolated from power to the second set of EPUs upon activation of the first pyrotechnical fuse. 4. The system described in clause H3, wherein the second high voltage bus includes a second pyrotechnical fuse, and wherein power to the third set of EPUs is isolated from power to the fourth set of EPUs upon activation of the second pyrotechnical fuse. 5. A system described in any one of clauses H1 to H4, wherein each battery includes a pyrotechnical fuse, and each pyrotechnical fuse is configured to disconnect the battery from the working portion of a corresponding one of the paired battery pack units. 6. The system of any one of clauses H1-H5, wherein the first paired battery pack unit high voltage wiring and the second paired battery pack unit high voltage wiring are electrically separate from each other. 7. A system as described in any one of clauses H1 to H6, wherein each EPU of the plurality of EPUs comprises a fuse for the high voltage power supply, and upon operation of the fuse, the associated EPU is disconnected from the operating parts of the aircraft. 8. The system of any one of clauses H1-H7, wherein each battery includes a charging contactor, and when the charging contactor opens, the associated battery is disconnected from the high voltage charging bus. 9. The system of any one of clauses H1-H8, wherein each set of EPUs comprises two EPUs. 10. The system of clause H9, wherein each set of EPUs comprises two diagonally symmetric EPUs. 11. The system described in clause H10, wherein the first paired battery pack unit is configured to provide power to the front outer EPU, the front inner EPU, the rear outer EPU, and the rear inner EPU. 12. The system described in clause H11, wherein the second paired battery pack unit is configured to provide power to the front outer EPU, the front inner EPU, the rear outer EPU, and the rear inner EPU. 13. The system described in any one of clauses H1 to H12, further comprising a third paired battery pack unit, the third paired battery pack unit comprising a fifth battery electrically connected to a sixth battery via a third high-voltage bus, the fifth battery configured to provide power to a fifth set of EPUs among the plurality of EPUs, and the sixth battery configured to provide power to a sixth set of EPUs among the plurality of EPUs, wherein the third high-voltage bus is electrically separate from the first high-voltage bus and the second high-voltage bus. 14. The system described in clause H13, wherein the fifth battery is configured to function as a backup battery for powering a sixth set of EPUs via the third high-voltage bus, and the sixth battery is configured to function as a backup battery for powering the fifth set of EPUs via the third high-voltage bus. 15. A power distribution system for an aircraft, comprising: a plurality of electric propeller units (EPUs); and a first paired battery pack unit, the first paired battery pack unit comprising 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, the first battery being configured to provide power to a first set of EPUs of the plurality of EPUs, the second battery being configured to provide power to a second set of EPUs of the plurality of EPUs, and the third battery being configured to provide power to a third set of EPUs of the plurality of EPUs. and a second paired battery pack unit, wherein the second paired battery pack unit comprises a fourth battery, a fifth battery, and a sixth battery, the fourth battery, the fifth battery, and the sixth battery are electrically connected via a second high-voltage bus, the fourth battery being configured to provide power to a fourth set of EPUs of the plurality of EPUs, the fifth battery being configured to provide power to a fifth set of EPUs of the plurality of EPUs, and the sixth battery being configured to provide power to a sixth set of EPUs of the plurality of EPUs, wherein the first high-voltage bus and the second high-voltage bus are electrically separate from each other. 16. A power distribution system for an aircraft, comprising: a plurality of electric propeller units (EPUs); and first paired battery pack units, the first paired battery pack units comprising 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, the first battery being configured to provide power to a first set of EPUs of the plurality of EPUs, the second battery being configured to provide power to a second set of EPUs of the plurality of EPUs, the third battery being configured to provide power to a third set of EPUs of the plurality of EPUs, and the fourth battery being configured to provide power to a fourth set of EPUs of the plurality of EPUs. and a second paired battery pack unit, wherein the second paired battery pack unit comprises a fifth battery, a sixth battery, a seventh battery, and an eighth battery, wherein the fifth battery, the sixth battery, the seventh battery, and the eighth battery are electrically connected via a second high-voltage bus, the fifth battery being configured to provide power to a fifth set of EPUs of the plurality of EPUs, the sixth battery being configured to provide power to a sixth set of EPUs of the plurality of EPUs, the seventh battery being configured to provide power to a seventh set of EPUs of the plurality of EPUs, and the eighth battery being configured to provide power to an eighth set of EPUs of the plurality of EPUs, wherein the first high-voltage bus and the second high-voltage bus are electrically separate from each other.

[0083] Clause Set I General Battery Management Systems 1. A system for battery management in a vehicle, comprising: 1. A system for battery management in a vehicle, comprising: a first battery pack, a second battery pack, a third battery pack; a first battery management system; a first paired battery pack unit comprising the first battery pack electrically connected to the second battery pack via a high-voltage bus, the first battery pack configured to power a first electric engine, and the second battery pack configured to power a second electric engine, the first paired battery pack unit being electrically separate from the third battery pack configured to power the third electric engine, the first battery pack configured to function as a backup battery pack for powering the second electric engine via the high-voltage bus, and the second battery pack configured to function as a backup battery for powering the first electric engine via the high-voltage bus, and the first battery management system detecting an electrical problem and blowing a fuse. 2. The system of clause I1, wherein the electrical problem is an overcurrent condition or a short circuit condition. 3. The system described in clause I2, 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 supply. 4. The system described in clause I2, wherein the electrical problem is associated with the high voltage bus and the blown fuse isolates the power supply to the first electric engine from the power supply to the second electric engine. 5. The system described in clause I2, wherein the electrical problem is associated with the first battery pack circuit and the blown fuse cuts off the first battery pack's supply of power to the first and second electric engines. 6. The system of any one of clauses I1-I5, wherein the first battery management system monitors the charge level for the first battery pack and sends information regarding the charge level to the charge control unit. 7. The system of clause I6, wherein the first battery management system monitors the first battery pack temperature and transmits information regarding the first battery pack temperature to the charge control unit. 8. The system described in clause I7, wherein the system further comprises a contactor, and the first battery management system opens the contactor and disconnects the first battery pack from the charging circuit upon receiving a signal from the charging control unit. 9. A system described in any one of clauses I1 to I8, wherein the fuse is a pyrotechnical fuse. 10. The system of any one of clauses I1 to I9, wherein the first electric engine and the second electric engine are powered aircraft electric propulsion units.

[0084] Clause Set J HV Charging Mode 1. A control system for charging an aircraft, the control system comprising: a battery pack; an input device configured to allow a user to select between different charging modes; two main contactors connecting the battery pack to an electric propulsion unit (EPU) load and an auxiliary load; an EPU load contactor connecting the battery pack to the EPU load; and a controller configured to receive a selected charging mode and control the contactors, wherein the controller is configured, upon receiving a user selection to charge in a first mode, to leave the two main contactors open and disconnect the EPU load and the auxiliary load; upon receiving a user selection to charge in a second mode, to close the two main contactors, leave the EPU load contactor open, connect the auxiliary load, and disconnect the EPU load; and upon receiving a user selection to charge in a third mode, to close the two main contactors and the EPU load contactor and connect the auxiliary load and the EPU load.

Claims

1. 1. A charging system for an aircraft, comprising: a plurality of electric propeller units (EPUs); a plurality of battery packs configured to power the plurality of EPUs; a charge control unit, the charge control unit comprising: determining a target charge level for each of the plurality of battery packs; receiving charge status information from each of the plurality of battery packs; 10. A charging system for an aircraft configured to command a battery pack of the plurality of battery packs to disconnect from charging upon determining that a target charge level for the battery pack has been reached.

2. 2. The system of claim 1, wherein the charge control unit is further configured to command a ground charging subsystem to cease charging upon determining that the target charge level for each of the plurality of battery packs has been reached.

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

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

5. The system of claim 4 , wherein the target charge level for at least one of the plurality of battery packs is determined to be different from the target charge level for another battery pack of the plurality of battery packs.

6. 6. The system of claim 1, wherein the charging control unit is further configured to receive fault status information from each of the plurality of battery packs and to instruct a battery pack of the plurality of battery packs to disconnect from charging upon determination of a fault associated with the battery pack.

7. The system of claim 6 , wherein the charging control unit is further configured to instruct a ground charging subsystem to cease charging upon determining the fault with the battery pack.

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

9. The system of 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 information, or battery pack temperature information.

10. 10. The system of claim 1, wherein the charging control unit is further configured to receive temperature status information from each of the plurality of battery packs and provide cooling instructions to a ground subsystem based on the temperature status information.

11. 1. A control unit for charging an aircraft, comprising: A charging control unit is provided, the charging control unit comprising: determining a target charge level for each of the plurality of battery packs; receiving charge status information from each of the plurality of battery packs; a control unit for charging the aircraft configured to command a battery pack of the plurality of battery packs to disconnect from charging upon determining that a target charge level for the battery pack has been reached.

12. 12. The control unit of claim 11, wherein the charging control unit is further configured to instruct a ground charging sub-control unit to cease charging the plurality of battery packs upon determining that the target charge level for each of the plurality of battery packs has been reached.

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

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

15. The control unit of claim 14 , wherein the target charge level for at least one of the plurality of battery packs is determined to be different from the target charge level for another battery pack of the plurality of battery packs.

16. The charging control unit receiving fault status information from each of the plurality of battery packs; 16. The control unit of claim 11, further configured to command a battery pack of the plurality of battery packs to disconnect from charging upon determination of a fault with the battery pack.

17. 17. The control unit of claim 16, wherein the charging control unit is further configured to instruct a ground charging sub-control unit to cease charging the battery pack upon determining the fault with the battery pack.

18. 18. The control unit of claim 17, wherein the charging control unit is further configured to provide an indication to the ground charging sub-control unit of which battery pack of the plurality of battery packs has failed.

19. The control unit of 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 information, or battery pack temperature information.

20. 20. The control unit of claim 11, wherein the charging control unit is further configured to receive temperature status information from each of the plurality of battery packs and provide cooling instructions to a ground sub-control unit based on the temperature status information.

21. 1. A method for charging an aircraft, comprising: determining, by one or more processors, a target charge level for each of the plurality of battery packs; receiving, by the one or more processors, charging status information from each of the plurality of battery packs; instructing, by the one or more processors, a battery pack of the plurality of battery packs to disconnect from charging upon determining that a target charge level for the battery pack has been reached.

22. 22. The method of claim 21, further comprising instructing, by the one or more processors, a ground charging sub-control unit to cease charging the plurality of battery packs upon determining that the target charge level for each of the plurality of battery packs has been reached.

23. 23. The method of claim 21 or 22, wherein determining the target charge level for each of the plurality of battery packs is based on flight information.

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

25. 25. The method of claim 24, wherein the target charge level for at least one of the plurality of battery packs is determined to be different from the target charge level for another battery pack of the plurality of battery packs.

26. 26. The method of any one of claims 21-25, further comprising: receiving, by the one or more processors, fault status information from a battery pack of the plurality of battery packs; and instructing, by the one or more processors, the battery pack to disconnect from charging upon determination of a fault associated with the battery pack.

27. 27. The method of claim 26, further comprising instructing, by the one or more processors, a ground charging sub-control unit to cease charging upon determining the fault with the battery pack.

28. 28. The method of claim 27, further comprising providing, by the one or more processors, to the ground charging sub-control unit an indication of which battery pack of the plurality of battery packs has failed.

29. The method of 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 information, or battery pack temperature information.

30. 30. The method of claim 21, 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, cooling instructions to a ground sub-control unit based on the temperature status information.

31. 1. An emergency responder detection system for an aircraft, comprising: at least one electric propeller 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; the battery management system is configured to receive aircraft motion information, detect that the motion information indicates a potential collision, detect a loss of current in the at least one low-voltage wire, and blow a battery pack fuse to disconnect the supply of high-voltage power.

32. 32. The system of claim 31, wherein the battery management system blows the battery pack fuse to disconnect the supply of high voltage power upon determining the potential crash occurred before the loss of current in the at least one low voltage wire.

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

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

35. 35. The system of claim 34, wherein the potential collision is detected based on the aircraft acceleration exceeding a threshold.

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

37. 37. The system of claim 36, wherein one of the battery management systems blows the battery pack fuses of all of the at least two battery packs upon determining the potential crash occurred before the loss of current in the at least one low-voltage wire.

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

39. at least two battery packs, each battery pack further comprising the battery management system and the connection to the at least one low voltage wire; 36. The system of claim 31, wherein the at least one low voltage wire for the at least two battery packs is bundled together and routed through a tail of the aircraft.

40. 40. The system of any one of claims 31 to 39, wherein the battery pack fuses are pyrotechnical fuses.

41. 1. A system for battery management for an aircraft, comprising: a battery management system including one or more processors, the one or more processors comprising: receiving aircraft movement information for the aircraft; detecting that the motion information indicates a potential collision; Detecting a loss of current in at least one low voltage wire; 1. A system for battery management for an aircraft, configured to blow a battery pack fuse of at least one battery pack configured to supply high-voltage power, thereby disconnecting the supply of high-voltage power by the at least one battery pack.

42. 42. The system of claim 41, wherein the battery management system blows the battery pack fuse to disconnect the supply of high voltage power upon determining the potential crash occurred before the loss of current in the at least one low voltage wire.

43. The battery management system includes: receiving a mode of the aircraft from a flight control system of the aircraft; determining a false collision detection based on detecting that the aircraft motion information indicates a potential collision, detecting that there is no loss of current in the at least one low-voltage wire, and detecting that the aircraft was in ground mode at the time of the potential collision; 43. The system of claim 41 or 42, further configured to receive new aircraft movement information after determining the false collision detection.

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

45. 45. The system of any of claims 44 to 44, wherein the potential collision is detected based on the aircraft acceleration exceeding a threshold.

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

47. 47. The system of claim 46, wherein one of the battery management systems blows the battery pack fuses of all of the at least two battery packs upon determining the potential crash occurred before the loss of current in the at least one low-voltage wire.

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

49. 46. ​​The system of any one of claims 41 to 45, further comprising at least two battery packs, each battery pack comprising the battery management system and a connection to the at least one low-voltage wire, the at least one low-voltage wire for the at least two battery packs being bundled together and routed through a tail section of the aircraft.

50. 50. The system of any one of claims 41 to 49, wherein the battery pack fuses are pyrotechnical fuses.

51. 1. A method for aircraft battery management, comprising: receiving, by the battery management system, aircraft motion information for the aircraft; detecting, by the battery management system, that the motion information indicates a potential collision; detecting, by the battery management system, a loss of current in at least one low voltage wire; and disconnecting, by the battery management system, a battery pack fuse of at least one battery pack configured to supply high-voltage power, thereby disconnecting the supply of high-voltage power by the at least one battery pack.

52. 52. The method of claim 51, further comprising, upon determination of the potential crash occurring prior to the loss of current in the at least one low voltage wire, blowing the battery pack fuse to disconnect the supply of high voltage power.

53. receiving, by the battery management system, a mode of the aircraft from a 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 loss of current in the at least one low voltage wire, and detecting that the aircraft was in ground mode at the time of the potential collision; 53. The method of claim 51 or 52, further comprising receiving, by the battery management system, new aircraft motion information after determining the false collision detection.

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

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

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

57. 57. The method of claim 56, wherein the at least one battery pack comprises a first battery pack and a second battery pack configured to supply the high-voltage power, and the battery management system blows all battery pack fuses of the first battery pack and the second battery pack upon determining the potential crash occurred before the loss of current in the at least one low-voltage wire.

58. 58. The method of any one of claims 51 to 57, wherein the battery pack fuses are pyrotechnical fuses.

59. 1. A charging system for an aircraft, comprising: a plurality of electric propeller units (EPUs); a plurality of battery packs configured to power the plurality of EPUs; a charging port configured to receive high voltage power for charging the plurality of battery packs; a common high voltage charging bus connected to the charging ports; the plurality of battery packs are charged through the common high voltage charging bus; 1. A charging system for an aircraft, wherein each of the plurality of battery packs includes a disconnecting device for disconnecting the battery pack from charging.

60. 60. The system of claim 59, wherein the common high voltage charging bus is electrically separate from high voltage wiring powering the plurality of EPUs.

61. 61. The system of claim 59 or 60, further comprising a high voltage channel for each battery pack, said high voltage channel connecting said battery packs to said common high voltage charging bus.

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

63. 63. The system of claim 62, wherein the disconnection device comprises a contactor.

64. 64. The system of claim 63, wherein the disconnection device comprises contactors on both the positive and negative sides of the high voltage channel.

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

66. 67. A system as claimed in any one of claims 59 to 66, wherein the charging port is located on a fuselage of the aircraft.

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

68. 68. The system of claim 67, wherein the charging port is further configured to accept communications from a charging control unit on board the aircraft.

69. 1. An electrical power distribution system for an aircraft, comprising: a plurality of electric propeller units (EPUs); a first paired battery pack unit, the first paired battery pack unit comprising a first battery electrically connected to a second battery via a first high voltage bus, the first battery configured to provide power to a first set of EPUs of the plurality of EPUs, and the second battery configured to provide power to a second set of EPUs of the plurality of EPUs; a second paired battery pack unit, the second paired battery pack unit comprising a third battery electrically connected to a fourth battery via a second high voltage bus, the third battery configured to provide power to a third set of EPUs of the plurality of EPUs, and the fourth battery configured to provide power to a fourth set of EPUs of the plurality of EPUs; 1. An electrical power distribution system for an aircraft, wherein the first high voltage bus and the second high voltage bus are electrically separate from each other.

70. the first battery is configured to function as a backup battery for powering the second set of EPUs via the first high voltage bus; the second battery is configured to function as a backup battery for powering the first set of EPUs via the first high-voltage bus; the third battery is configured to function as a backup battery for powering the fourth set of EPUs via the second high voltage bus; 70. The system of claim 69, wherein the fourth battery is configured to function as a backup battery for powering the third set of EPUs via the second high voltage bus.

71. 71. The system of claim 69 or 70, wherein the first high voltage bus comprises a first pyrotechnical fuse, and wherein the power to the first set of EPUs is separate from the power to the second set of EPUs upon activation of the first pyrotechnical fuse.

72. 72. The system of claim 71, wherein the second high voltage bus comprises a second pyrotechnical fuse, and wherein the power to the third set of EPUs is separate from the power to the fourth set of EPUs upon activation of the second pyrotechnical fuse.

73. 73. The system of any one of claims 69 to 72, wherein each battery comprises a pyrotechnical fuse, each pyrotechnical fuse configured to disconnect the battery from the working portion of a corresponding one of the paired battery pack units.

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

75. 75. A system as described in any one of claims 69 to 74, wherein each EPU of the plurality of EPUs comprises a fuse for a high voltage power supply, and upon operation of the fuse the associated EPU is disconnected from the operating portion of the aircraft.

76. A system as claimed in any one of claims 69 to 75, wherein each battery comprises a charging contactor, and when the charging contactor is opened, the associated battery is disconnected from the high voltage charging bus.

77. 77. The system of any one of claims 69 to 76, wherein each set of the sets of EPUs comprises two EPUs.

78. 78. The system of claim 77, wherein each set of the sets of EPUs comprises two diagonally symmetric EPUs.

79. 79. The system of claim 78, wherein the first paired battery pack unit is configured to provide power to a front outer EPU, a front inner EPU, a rear outer EPU, and a rear inner EPU.

80. 80. The system of claim 79, wherein the second paired battery pack unit is configured to provide power to the front outer EPU, the front inner EPU, the rear outer EPU, and the rear inner EPU.

81. a third paired battery pack unit, the third paired battery pack unit comprising a fifth battery electrically connected to a sixth battery via a third high voltage bus, the fifth battery configured to provide power to a fifth set of EPUs of the plurality of EPUs, and the sixth battery configured to provide power to a sixth set of EPUs of the plurality of EPUs; 81. The system of any one of claims 69 to 80, wherein the third high voltage bus is electrically separate from the first high voltage bus and the second high voltage bus.

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

83. 1. An electrical power distribution system for an aircraft, comprising: a plurality of electric propeller units (EPUs); a first paired battery pack unit, the first paired battery pack unit comprising 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, the first battery being configured to provide power to a first set of EPUs of the plurality of EPUs, the second battery being configured to provide power to a second set of EPUs of the plurality of EPUs, and the third battery being configured to provide power to a third set of EPUs of the plurality of EPUs; a second paired battery pack unit, the second paired battery pack unit comprising 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, the fourth battery being configured to provide power to a fourth set of EPUs of the plurality of EPUs, the fifth battery being configured to provide power to a fifth set of EPUs of the plurality of EPUs, and the sixth battery being configured to provide power to a sixth set of EPUs of the plurality of EPUs; 1. An electrical power distribution system for an aircraft, wherein the first high voltage bus and the second high voltage bus are electrically separate from each other.

84. 1. An electrical power distribution system for an aircraft, comprising: a plurality of electric propeller units (EPUs); a first paired battery pack unit, the first paired battery pack unit comprising 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, the first battery being configured to provide power to a first set of EPUs of the plurality of EPUs, the second battery being configured to provide power to a second set of EPUs of the plurality of EPUs, the third battery being configured to provide power to a third set of EPUs of the plurality of EPUs, and the fourth battery being configured to provide power to a fourth set of EPUs of the plurality of EPUs; a second paired battery pack unit, the second paired battery pack unit comprising 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, the fifth battery being configured to provide power to a fifth set of EPUs of the plurality of EPUs, the sixth battery being configured to provide power to a sixth set of EPUs of the plurality of EPUs, the seventh battery being configured to provide power to a seventh set of EPUs of the plurality of EPUs, and the eighth battery being configured to provide power to an eighth set of EPUs of the plurality of EPUs; 1. An electrical power distribution system for an aircraft, wherein the first high voltage bus and the second high voltage bus are electrically separate from each other.

85. 1. A system for battery management in a vehicle, comprising: a first battery pack; a second battery pack; a third battery pack; and a first battery management system; a first paired battery pack unit comprising the first battery pack electrically connected to the second battery pack via a high voltage bus, the first battery pack configured to power a first electric engine and the second battery pack configured to power a second electric engine; the first paired battery pack unit being electrically separate from the third battery pack configured to power a third electric engine; the first battery pack is configured to function as a backup battery pack to power the second electric engine via the high voltage bus; the second battery pack is configured to function as a backup battery for powering the first electric engine via the high voltage bus; A system for battery management in a vehicle, wherein the first battery management system detects an electrical problem and blows a fuse.

86. 86. The system of claim 85, wherein the electrical problem is an overcurrent condition or a short circuit condition.

87. 87. The system of 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 an electrical power supply.

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

89. 87. The system of claim 86, wherein the electrical problem is associated with a first battery pack circuit, and the blowing of the fuse isolates the first battery pack from providing electrical power to the first electric engine and a second electric engine.

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

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

92. 92. The system of claim 91, further comprising a contactor, wherein the first battery management system opens the contactor to disconnect the first battery pack from a charging circuit upon receiving a signal from the charging control unit.

93. A system according to any one of claims 85 to 92, wherein the fuse is a pyrotechnical fuse.

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

95. 1. A control system for charging an aircraft, comprising: A battery pack; an input device configured to allow a user to select between different charging modes; two main contactors connecting the battery pack to an electric propulsion unit (EPU) load and an auxiliary load; an EPU load contactor connecting the battery pack to the EPU load; a controller configured to receive the selected charging mode and to control the contactors; the controller is configured to, upon receiving a user selection to charge in a first mode, leave the two main contactors open and disconnect the EPU load and the auxiliary load; the controller is configured, upon receiving a user selection to charge in a second mode, to close the two main contactors, leave the EPU load contactor open, connect the auxiliary load, and disconnect the EPU load; A control system for charging an aircraft, wherein the controller is configured, upon receiving a user selection to charge in a third mode, to close the two main contactors and the EPU load contactor and connect the auxiliary load and the EPU load.

Images