Patent Description:
Electric aircrafts include battery packs to power various flight components, including electric propulsion units (EPUs) which enable flight. These battery packs are critical to ensure the EPUs can provide the aircraft's lift and thrust support. Therefore, there is a need to provide for redundancy in the aircraft's power system to avoid a single point of failure. There is also a need to ensure a fault or failure condition does not propagate and damage other critical aircraft components. The disclosed high voltage power system solves these problems and other problems by connecting battery packs together in a battery pack unit, where each battery pack in a unit acts as a backup for the others. Further, each battery back unit is electrically separate from other battery pack units.

Additionally, to ensure battery packs can power the EPUs for the duration of a flight, battery packs need to be sufficiently charged prior to take-off. Therefore, there is a need to charge battery packs efficiently and effectively. The disclosed high voltage power system solves these problems and others by controlling battery pack charge amount based on upcoming flight information, historical battery pack information, and a monitored state of the battery pack. The disclosed high voltage power system also solves this problem and others by providing a single point of charging for multiple battery packs.

Finally, in the event of a crash, there is a need for a first responder to be able to shut off the high voltage power system quickly and safely. The disclosed high voltage power system solves this problem by providing low voltage cut loops connected to the battery packs. Upon detecting that a first responder has cut a low voltage cut loop, fuses to the battery packs are blown and the high voltage power system is no longer powered. The cut loop may be routed to the tail of the aircraft to provide separation from high voltage lines and increase safety for the first responder.

<CIT> relates to electrically-driven propulsors for regional aircraft. These combine fans optimized for quiet regional operations with rotating variable pitch mechanisms contained within the hubs and highly efficient ring electric motors. The ring motors are used to drive annular fans at the inner or outer radius, or a conventional fan at the inner radius. The assembly combines a single fan, fixed or variable pitch, with a row of stators, or multiple counter-rotating fans without stators.

The present disclosure generally relates to a power system for an aircraft. One aspect of the present disclosure provides a method for aircraft battery management. The method comprises receiving, by a battery management system, aircraft movement information of an aircraft; detecting, by the battery management system, that the movement information indicates a potential crash; 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 supply of the high voltage power by the at least one battery pack. Another aspect of the present disclosure provides a system for an aircraft battery management. The system comprises a battery management system (<NUM>) including one or more processors; wherein the one or more processors are configured to execute instructions to cause the battery management system to perform the above method.

The present disclosure addresses components of electric vertical takeoff and landing (eVTOL) aircraft primarily for use in a non-conventional aircraft. For example, the eVTOL aircraft of the present disclosure may be intended for frequent (e.g., over <NUM> flights per workday), short-duration flights (e.g., less than <NUM> miles per flight) over, into, and out of densely populated regions. The aircraft may be intended to carry <NUM>-<NUM> passengers or commuters who have an expectation of a low-noise and low-vibration experience. Accordingly, it may be desired that their components are configured and designed to withstand frequent use without wearing, that they generate less heat and vibration, and that the aircraft include mechanisms to effectively control and manage heat or vibration generated by the components. Further, it may be intended that several of these aircraft operate near each other over a crowded metropolitan area. Accordingly, it may be desired that their components are configured and designed to generate low levels of noise interior and exterior to the aircraft, and to have a variety of safety and backup mechanisms. For example, it may be desired for safety reasons that the aircraft are propelled by a distributed propulsion system, avoiding the risk of a single point of failure, and that they are capable of conventional takeoff and landing on a runway. Moreover, it may be desired that the aircraft can safely vertically takeoff and land from and into relatively restricted spaces (e.g., vertiports, parking lots, or driveways) compared to traditional airport runways while transporting around <NUM>-<NUM> passengers or commuters with accompanying baggage. These use requirements may place design constraints on aircraft size, weight, operating efficiency (e.g., drag, energy use), which may impact the design and configuration of the aircraft components.

Disclosed embodiments provide new and improved configurations of aircraft components that are not observed in conventional aircraft, and/or identified design criteria for components that differ from those of conventional aircraft. Such alternate configurations and design criteria, in combination addressing drawbacks and challenges with conventional components, yielded the embodiments disclosed herein for various configurations and designs of eVTOL aircraft components.

In some embodiments, the eVTOL aircraft of the present disclosure may be designed to be capable of both vertical and conventional takeoff and landing, with a distributed electrical propulsion system enabling vertical flight, forward flight, and transition. Thrust may be generated by supplying high voltage electrical power to the electric engines of the distributed electrical propulsion system, which each may convert the high voltage electrical power into mechanical shaft power to rotate a propeller. Embodiments disclosed herein may involve optimizing the energy density of the electrical propulsion system. Embodiments may include an electric engine connected to an onboard electrical power source, which may include a device capable of storing energy such as a battery or capacitor, or may include one or more systems for harnessing or generating electricity such as a fuel powered generator or solar panel array. Some disclosed embodiments provide for weight reduction and space reduction of components in the aircraft, thereby increasing aircraft efficiency and performance. Given focus on safety in passenger transportation, disclosed embodiments implement new and improved safety protocols and system redundancy in the case of a failure, to minimize any single points of failure in the aircraft propulsion system. Some disclosed embodiments also provide new and improved approaches to satisfying aviation and transportation laws and regulations.

<FIG> illustrates an example eVTOL aircraft, consistent with embodiments of the present disclosure. As shown in <FIG>, in some embodiments, the distributed electrical propulsion system of the eVTOL aircraft <NUM> may include twelve electric engines <NUM>, which may be mounted on booms forward and aft of the main wings of the aircraft <NUM>. The forward electric engines <NUM> may be tiltable mid-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 engines <NUM> may be of a clockwise type or counterclockwise type in terms of direction of propeller rotation. The aft electric engines <NUM> may be fixed in a vertically oriented position (e.g., to generate vertical lift), and may also be of a clockwise type or counterclockwise type in terms of direction of propeller rotation.

The aircraft <NUM> may possess various combinations of forward and aft electric engines <NUM>. For example, in some embodiments, the aircraft <NUM> may possess six forward electric engines <NUM> and six aft electric engines <NUM>. In some other embodiments, the aircraft <NUM> may include four forward electric engines <NUM> and four aft electric engines <NUM>, or any other combination of forward and aft engines <NUM>. In some other embodiments, the number of forward electric engines and aft electric engines are not equivalent.

In some embodiments, for a vertical takeoff and landing (VTOL) mission, the forward electric engines <NUM> as well as aft electric engines <NUM> may provide vertical thrust during takeoff and landing. During flight phases where the aircraft <NUM> is in forward flight-mode, the forward electric engines <NUM> may provide horizontal thrust, while the propellers of the aft electric engines <NUM> may be stowed at a fixed position in order to minimize drag. The aft electric engines <NUM> may be actively stowed with position monitoring.

In some embodiments, in a conventional takeoff and landing (CTOL) mission, the forward electric engines <NUM> may provide horizontal thrust for wing-borne take-off, cruise, and landing. In some embodiments, the aft electric engines <NUM> may not be used for generating thrust during a CTOL mission and the aft propellers may be stowed in place.

Transition from vertical flight to forward flight and vice-versa may be accomplished via the tilt propeller subsystem. The tilt propeller subsystem may redirect thrust between a primarily vertical direction during vertical flight mode to a mostly horizontal direction during forward-flight mode. A variable pitch mechanism may change the forward electric engine's propeller-hub assembly blade collective angles for operation during the hover-phase, transition phase, and cruise-phase.

The tilt propeller system may include a linear or rotary actuator to change the orientation of a 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, a rotary actuator may include a motor, inverter, and gearbox. In some embodiments, a gearbox may include various types of gears interfacing to provide a gear reduction capable of orienting the propulsion system. In some embodiments, a tilt propeller system may include a redundant configuration such that multiple motors, inverters, and gearboxes are present and interface using a gear. In some embodiments, a configuration utilizing multiple motors, gearboxes, and inverters may allow a failed portion of the redundant configuration to be driven by the motor, inverter, and gearbox of another portion of the configuration. In some embodiments, a gearbox configuration may also allow the tilt propeller system to maintain a propulsion system orientation with the help of, or without, additional power being provided by the system.

In some embodiments, an electric engine <NUM> may be housed or connected to a boom of the aircraft <NUM> and include a motor, inverter, and gearbox. In some embodiments, the motor, inverter, and gearbox may be interfaced such that they share a central axis. In some embodiments, the torque originating in the motor may be sent away from the propellers of the propulsion system and to a gearbox. In some embodiments, a gearbox may provide a gear reduction and then send the torque, via a main shaft, back through a bearing located inside the motor and to the propeller. In some embodiments, an inverter may be mounted on the rear of a gearbox such that a main shaft does not travel through the inverter when outputting torque to the propeller.

As shown in <FIG>, the aircraft <NUM> may be configured with a distributed electric propulsion system enabling vertical flight, forward flight, and transition. The forward <NUM> electric engines <NUM> (which are numbered <NUM>-<NUM> from left to right) are with variable pitch propellers tilt to achieve vertical takeoff and landing, transition flight and fully wing-borne flight. The aft <NUM> electric engines <NUM> (which are numbered <NUM>-<NUM> from left to right) are equipped with fixed pitch propellers that operate during vertical takeoff and landing and transition and are stowed in a minimum drag position for conventional flight. The flight controls are an integrated fly-by-wire system that features envelope protection and structural load limiting functions. The aircraft <NUM> will be equipped with advanced cockpit avionics, a flight management system, and the sensors necessary to support the intended operations and system functions.

In some embodiments, an electrical propulsion system (EPS) as described herein may generate thrust by supplying High Voltage (HV) electric power to the electric engine <NUM>, which in turn converts HV power into mechanical shaft power which is used to rotate a propeller. As mentioned above, an aircraft <NUM> as described herein may possess multiple electric engines <NUM> which are boom-mounted forward and aft of the wing. The amount of thrust each electric engine <NUM> generates may be governed by a torque command from the Flight Control System (FCS) over a digital communication interface to each electric engine <NUM>. Embodiments may include forward electric engines <NUM>, and may be able to alter their orientation, or tilt. Additional embodiments include forward engines that may be a clockwise (CW) type 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.

In some embodiments, the aircraft <NUM> includes a high voltage power supply (HVPS) system to supply the High Voltage (HV) electric power. The HVPS system is the source of power on the aircraft <NUM> and configured to distribute the stored electrical energy to other systems on the aircraft <NUM>, including the electrical propulsion system (EPS) for converting electrical power into mechanical rotational shaft power to generate thrust. As shown in <FIG>, the HVPS system of the aircraft <NUM> may include six battery packs <NUM> (which are numbered <NUM>-<NUM> from left to right) installed within the battery bays in the wing of the aircraft <NUM>. In some embodiments, six battery packs <NUM> may have the identical design, to simplify the design, manufacturing, and logistics. The battery packs <NUM> may power one or more electric engines <NUM>. While six battery packs <NUM> are shown, the aircraft <NUM> may have any number of battery packs <NUM>.

In some embodiments, a single battery pack <NUM> may be electrically connected to, and power, multiple electric engines <NUM>. For example, in some embodiments, a battery pack <NUM> may power an electric engine <NUM> on either side of a longitudinal axis. In some embodiments a battery pack <NUM> may power an electric engine <NUM> on either side of a horizontal axis. In some embodiments, as shown in <FIG>, a battery pack <NUM> may power two diagonally opposing electric engines <NUM>. For example, battery pack <NUM> may power electric engines <NUM> and <NUM>. Battery pack <NUM> may power electric engines <NUM> and <NUM>. Battery pack <NUM> may power electric engines <NUM> and <NUM>. Battery pack <NUM> may power electric engines <NUM> and <NUM>. Battery pack <NUM> may power electric engines <NUM> and <NUM>. Battery pack <NUM> may power electric engines <NUM> and <NUM>. Therefore, upon a loss of a battery pack <NUM>, the impact to roll or pitch moments can be reduced because the loss of lift is balanced. In some embodiments, battery packs <NUM> may power different arrangements of electric engines <NUM> to reduce roll, pitch, or yaw moments that may be caused by a loss of the battery pack <NUM>. For example, in some embodiments, battery packs <NUM> may be connected to electric engines <NUM> in any manner that balances lift and/or thrust across the longitudinal and horizontal axis of the aircraft.

Further, the HVPS system includes a cross-link <NUM> possessing at least one fuse allowing for pairing of two or more battery packs <NUM>. Through the cross-link, power for the electric engines <NUM> can be shared among the paired battery packs <NUM>. Therefore, multiple battery packs <NUM> can simultaneously power multiple electric engines <NUM>. This arrangement provides for redundancy and avoids a single point of failure because each paired battery <NUM> may act as a backup for the other(s). Upon failure of a battery pack <NUM>, one or more connected battery packs <NUM> may continue powering the failed battery pack's connected electric engines <NUM>.

In some embodiments, as shown in <FIG>, a pair of battery packs <NUM> may include two battery packs <NUM>. In some embodiments, a pair of two battery backs <NUM> may power a total of four electric engines <NUM>. For example, battery pack <NUM>, providing power to electric engines <NUM> and <NUM>, may be cross-linked to battery pack <NUM>, providing power to electric engines <NUM> and <NUM>. Battery pack <NUM>, providing power to electric engines <NUM> and <NUM>, may be cross linked to battery pack <NUM>, providing power to electric engines <NUM> and <NUM>. Battery pack <NUM>, providing power to electric engines <NUM> and <NUM>, may be cross linked to battery pack <NUM>, providing power to electric engines <NUM> and <NUM>.

<FIG> illustrates another example eVTOL aircraft, consistent with embodiments of the present disclosure. In some embodiments, electric engines <NUM> may include multiple motor stages that are each independently powered by different battery packs <NUM> so that should one battery pack <NUM> fail only a portion of the EPU is unpowered and the EPU can continue operating at a reduced power level. In some embodiments an electric engine <NUM> may include two partial motors. For example, battery pack <NUM> may power first partial motors on electric engines <NUM>, <NUM>, <NUM>, and <NUM>. Battery pack <NUM> may power second partial motors on electric engines <NUM>, <NUM>, <NUM>, and <NUM>. In some embodiments, different configurations may be used. For example, two battery packs may provide power to partial motors on electric engines <NUM>, <NUM>, <NUM>, and <NUM>.

<FIG> illustrates an electric engine <NUM> with 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 <NUM>. The two partial motors 191a and 191b can operate independently to drive blades of an EPU and can operate simultaneously to drive the blades at a 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 so that each electric engine <NUM> has an electrically separate backup.

The above configurations are provided as an example, but a different number and configuration of battery packs <NUM>, electric engines <NUM>, battery pack to electric engine connections, and battery pack cross link combinations may be used. In some embodiments, each battery battery pack <NUM> may power an individual electric engine <NUM>. For example, an aircraft may have four, six, eight, ten, twelve, or any number of electric engines <NUM> and the number of battery packs <NUM> may match the number of electric engines. In some embodiments, each battery pack <NUM> may power only one electric engine <NUM> and may be electrically separate from all other battery packs <NUM>. In some embodiments, each battery pack <NUM> may power one or more partial motors and each electric engine may include two or more partial motors. Therefore, each electric engine <NUM> may have a backup power source but the battery packs <NUM> are still electrically separate.

In some embodiments, each battery pack <NUM> may power multiple electric engines <NUM>. As described above, battery packs <NUM> may power sets of electric engines <NUM> that are symmetrical across one or more axes of symmetry. In some embodiments, a battery pack <NUM> may power electric engines <NUM> that are symmetrical across an aircraft's longitudinal axis, lateral axis, or both. For example, as described above, in some embodiments, different battery packs <NUM> may power diagonally symmetric electric engines <NUM> and <NUM>, <NUM> and <NUM>, <NUM> and <NUM>, <NUM> and <NUM>, <NUM> and <NUM>, and <NUM> and <NUM>.

In some embodiments, a battery pack <NUM> may power more than two electric engines <NUM>. In some embodiments, a battery pack <NUM> may power two or more sets of diagonally symmetric electric engines. For example, in some embodiments, a battery pack <NUM> may power electric engines <NUM>, <NUM>, <NUM>, and <NUM>, where electric engines <NUM> and <NUM> are diagonally symmetric and electric engines <NUM> and <NUM> are diagonally symmetric. In some embodiments, the set of electric engines <NUM> powered by a battery pack <NUM> may include an inboard diagonally symmetric pair of electric engines <NUM> and an outboard diagonally symmetric pair of electric engines <NUM>.

In some embodiments, a battery pack <NUM> may power four or more electric engines <NUM> in a configuration that is symmetric across the longitudinal axis of symmetry. For example, battery pack <NUM> may power electric engines <NUM>, <NUM>, <NUM>, and <NUM>. In some embodiments, in each of the above configurations, a battery pack <NUM> may provide power to one or more partial motors and each electric engine <NUM> may include two or more partial motors. Therefore, each electric engine <NUM> may have a backup power source but the battery packs <NUM> are still electrically separate.

In some embodiments, some or all of the battery packs <NUM> are interconnected. As described above, a cross-link <NUM> may allow each battery pack <NUM> to act as backup power for another. For example, in some embodiments, battery pack <NUM> may directly power a first number of electric engines and a second battery pack <NUM> may directly power a second number of electric engines. The first and second battery packs <NUM> may be cross-linked together to form a battery pack unit. Therefore, each battery pack in the unit may act as a backup for the other. Upon failure of a battery pack in the unit, the failing battery pack may be disconnected and electric engines <NUM> will be powered by one or more non-failing battery packs in the unit. The battery packs in a battery pack unit may be electrically separate from other battery pack units.

As described above, in some embodiments, a battery pack unit may comprise two battery packs <NUM>, wherein each battery pack <NUM> powers a number of electric engines <NUM>. As described above, in some embodiments, each battery pack <NUM> may power two diagonally symmetric electric engines <NUM>. Therefore, each battery pack unit may power a total of four electric engines <NUM> and each electric engine has a battery pack backup. In some embodiments, each battery pack <NUM> in a battery pack unit may power four electric engines <NUM>, comprising two sets of diagonally symmetric electric engines <NUM>. Therefore, each battery pack unit may power a total of eight electric engines <NUM> and each electric engine has a battery pack backup.

In some embodiments, a battery pack unit may comprise three battery packs <NUM>, wherein each battery pack powers a number of electric engines <NUM>. For example, in some embodiments, each battery pack <NUM> may power two diagonally symmetric electric engines <NUM>. Therefore, each battery pack unit may power a total of six electric engines <NUM> and each electric engine <NUM> has two battery pack backups. In some embodiments, each battery pack <NUM> in the battery pack unit may power four electric engines <NUM>, comprising two sets of diagonally symmetric electric engines <NUM>. Therefore, each battery pack unit may power a total of twelve electric engines <NUM> and each electric engine has two battery pack backups.

In some embodiments, a battery pack unit may comprise four battery packs <NUM>, wherein each battery pack powers a number of electric engines. For example, in some embodiments, each battery pack may power two diagonally symmetric electric engines <NUM>. Therefore, each battery pack unit may power a total of eight electric engines <NUM> and each electric engine has three battery pack backups. In other embodiments, each battery pack <NUM> in the battery pack unit may power four electric engines <NUM>, comprising two sets of diagonally symmetric electric engines <NUM>. Therefore, each battery pack unit may power sixteen electric engines <NUM> and each electric engine has three battery pack backups.

In some embodiments, electric engines <NUM> comprise a single motor that is powered by the one or more battery packs <NUM>. In some embodiments, each electric engine <NUM> may include two or more partial motors and the battery packs <NUM> may power partial motors. In some embodiments, the electric engine <NUM> powering configurations described above may include powering a partial motor of a battery pack. For example, in some embodiments, each electric engine <NUM> may include two partial motors and a battery pack unit may power partial motors of electric engines <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, and <NUM>. A second battery pack unit may power partial motors of electric engines <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, and <NUM>. A third battery pack unit may power partial motors of electric engines <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, and <NUM>. A fourth battery pack unit may power partial motors of electric engines <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, and <NUM>. Therefore, each electric engine <NUM> will receive backup power through the other partial motor. As described above, each battery pack unit may comprise one or more battery packs. For example, a battery pack unit may comprise one, two, three, or four battery packs.

In some embodiments, each electric engine may include two partial motors and a battery pack unit may power partial motors of electric engines <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, and <NUM>. A second battery pack unit may power partial motors of electric engines <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, and <NUM>. A third battery pack unit may power partial motors of electric engines <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, and <NUM>. Therefore, each electric engine <NUM> will receive backup power through the other partial motor. As described above, each battery pack unit may comprise one or more battery packs. For example, a battery pack unit may comprise one, two, three, or four battery packs.

In some embodiments, each electric engine may include two partial motors and a battery pack unit may power partial motors of electric engines <NUM>, <NUM>, <NUM>, and <NUM>. A second battery pack unit may power partial motors of electric engines <NUM>, <NUM>, <NUM>, and <NUM>. A third battery pack unit may power partial motors of electric engines <NUM>, <NUM>, <NUM>, and <NUM>. A fourth battery pack unit may power partial motors of electric engines <NUM>, <NUM>, <NUM>, and <NUM>. A fifth battery pack unit may power partial motors of electric engines <NUM>, <NUM>, <NUM>, and <NUM>. A sixth battery pack unit may power partial motors of electric engines <NUM>, <NUM>, <NUM>, and <NUM>. Therefore, each electric engine <NUM> will receive backup power through the other partial motor. As described above, each battery pack unit may comprise one or more battery packs. For example, a battery pack unit may comprise one, two, three, or four battery packs. Different configurations of battery packs <NUM>, electric engines <NUM>, battery pack to electric engine connections, and battery pack cross link combinations may be chosen to best balance aircraft power needs, system redundancy, and fault tolerance.

<FIG> illustrates a diagram of a high voltage power system for an eVTOL aircraft, consistent with embodiments of the present disclosure. As shown in <FIG>, eVTOL aircraft may include a battery assembly comprising electrically separate battery pack units (e.g. <NUM>, <NUM>, and <NUM>). Each battery pack unit may include battery packs <NUM> that are cross-linked together, as described above. In some embodiments, the battery pack units may include battery packs <NUM> that ensure aircraft controllability is maintained upon the loss of a battery pack unit. Therefore, upon a 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 <NUM> that power electric engines <NUM> on opposite sides of one or more axis of symmetry. Therefore, upon a loss of a battery pack unit, the impact to roll, pitch, or yaw moments can be reduced because the loss of lift and/or thrust is balanced. In some embodiments, loss of power, or reduction of power, caused by failure of a battery pack unit will have a substantially symmetric effect (e.g., <±<NUM>%, <±<NUM>%, <±<NUM>%, <±<NUM>%, or <±<NUM>% asymmetry) with respect to roll, pitch, and/or yaw of the aircraft. In some embodiments, the battery pack units may include battery packs <NUM> to reduce an overall amount of high voltage wiring between the battery packs. In some embodiments, the battery pack units may include battery packs <NUM> to minimize power requirements.

In some embodiments, as shown in <FIG>, the HVPS system may comprise three electrically separate battery pack units. For example, in some embodiments, battery pack unit <NUM> may include battery packs <NUM> and <NUM>, powering electric engines <NUM>, <NUM>, <NUM>, and <NUM>. Battery pack unit <NUM> may include battery packs <NUM> and <NUM>, powering electric engines <NUM>, <NUM>, <NUM>, and <NUM>. Battery pack unit <NUM> may include battery packs <NUM> and <NUM>, powering electric engines <NUM>, <NUM>, <NUM>, and <NUM>. Therefore, each battery pack unit may include two paired battery packs <NUM> that simultaneously power four electric engines <NUM>. Upon the failure of one battery pack <NUM> in a battery pack unit, the other paired battery pack <NUM> will continue powering the four electric engines.

In some embodiments, each battery pack units <NUM>, <NUM>, <NUM> may include a high voltage bus to cross-link battery packs <NUM> within the battery pack unit. In some embodiments, the cross-link <NUM> connects two high voltage channels, each feeding one or more electric engines <NUM>. For example, in some embodiments, the cross link <NUM> may be connected to each battery pack's high voltage channel before the channel splits to power multiple electric engines <NUM> (e.g. to power two electric engines). A cross link may further include a bus connecting the negative voltage channels after the negative voltage channels are combined (e.g. after powering two electric engines).

In some embodiments, each cross link <NUM> may include at least one fuse to disconnect the cross-link upon a failure of the cross-link. For example, fuses <NUM>, <NUM>, and <NUM> may be located on the cross-link connection of the positive high voltage channels in battery pack units <NUM>, <NUM>, and <NUM>. In some embodiments, the fuses may be pyro-technical fuses. As further detailed below, a battery management system of a connected battery pack <NUM> may determine a failure in a cross-link, such as a short circuit or overcurrent condition, and blow the associated pyro-technical fuse. Therefore, the cross-link can be disconnected and further damage to HVPS system components (e.g. electric engines, batteries, EPUS) can be avoided. Further, the electric engines <NUM> will still receive power from the paired battery pack <NUM> in the battery pack unit. For example, upon a cross-link failure, pyro-technical fuse <NUM> may be blown, but electric engines <NUM> and <NUM> will still receive power from battery pack <NUM>, and electric engines <NUM> and <NUM> will still receive power from battery pack <NUM>.

In some embodiments, there may be additional pyrotechnical fuses on the cross-link connection of the negative high voltage channels. For example, pyrotechnical fuses <NUM>, <NUM>, and <NUM> may be located on the cross-links in battery pack units <NUM>, <NUM>, and <NUM>, respectively. This configuration may provide additional redundancy for 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 positive cross link connection does not blow after being commanded to, a connected battery management system can instruct the negative cross link fuse to blow. Further, 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.

In some embodiments, the HVPS system may include load disconnection devices to disconnect a portion of the HVPS circuit upon a failure (e.g. short circuit or overcurrent condition) of a downstream electric engine, a downstream EPU, or other downstream distribution circuitry. In some embodiments, a load disconnection device may be located directly upstream of the electric engine. For example, in some embodiments, load disconnection devices <NUM>, <NUM>, <NUM>, and <NUM> may be located on the high voltage channel powering engines <NUM>, <NUM>, <NUM>, and <NUM>, respectively. Load disconnection devices <NUM>, <NUM>, <NUM>, and <NUM> may be located on the high voltage channels powering engines <NUM>, <NUM>, <NUM>, and <NUM>, respectively. Load disconnection devices <NUM>, <NUM>, <NUM>, and <NUM> may be located on the high voltage channels powering engines <NUM>, <NUM>, <NUM>, and <NUM>, respectively.

In some embodiments, the load disconnection devices are pyro-technical fuses. Upon failure of a downstream component, the pyro-technical fuse may receive a signal (e.g. from a battery management system of a connected battery) and blow the fuse. Therefore, the downstream components can be disconnected and further damage to other equipment (e.g. electric engines, batteries, EPUS) can be avoided. Further, the remaining electric engines <NUM> in the battery pack unit will still receive power from the connected battery packs <NUM>. For example, upon a failure in a device or wiring downstream of pyrotechnical fuse <NUM>, the pyrotechnical fuse <NUM> may be blown, but electric engines <NUM>, <NUM>, and <NUM> will still receive power from battery packs <NUM> and <NUM>. Further, in some embodiments, the load disconnection device may include a contactor and the battery management system may command the contactor to disconnect the circuit. In some embodiments, both a contactor and a fuse may be used to provide for additional redundancy and the pyro-technical fuse may act as a backup for the contactor.

In some embodiments, the HVPS system may include a high voltage charging channel allowing all the battery packs <NUM> to be charged from the same charging port. The high voltage charging channel may include charging disconnection devices. In some embodiments, the charging disconnection devices may be positioned downstream of a common charging bus on the positive charging side. For example, disconnection devices <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, and <NUM> may provide for disconnection of battery packs <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, and <NUM>, respectively. Similarly, in some embodiments, additional charging disconnection devices may be positioned upstream of a common charging bus on the negative charging side. For example, disconnection devices <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, and <NUM> may provide for disconnection of battery packs <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, and <NUM>, respectively.

In some embodiments, the charging disconnection devices are contactors, such as K4 Pos and K4 Neg in <FIG>. The charging contactors may act as a redundant measure to disconnect the battery packs <NUM> from charging. As further detailed below, a battery pack <NUM> may report a charging issue to a charge control unit (CCU). For example, a battery pack <NUM> may report a short circuit or overcurrent condition in the battery pack <NUM> or in the high voltage charging channel. In some embodiments, if the CCU fails to stop the charging, the battery packs <NUM> may command the charging contactors to disconnect the charging channel. In some embodiments, the battery packs <NUM> may automatically command the charging contactors to disconnect the charging channel without waiting for the CCU to fail. In some embodiments, after commanding the CCU to stop charging and/or disconnecting a battery pack <NUM> that detected the charging issues, the battery pack <NUM> and/or CCU may command the other battery packs <NUM> to disconnection from the charging channel. By disconnecting the battery packs <NUM> upon detecting a charging issue, damage to HVPS components can be avoided.

<FIG> illustrates a circuit diagram for a High Voltage Junction Box (HVJB), consistent with embodiments of the present disclosure. HVJB <NUM> may be electrically connected to the HV loads <NUM> to provide high voltage power. Specifically, the DC/DC converter in the battery management system (BMS) and the power storage element BT1 (e.g., the battery cells connected in parallel and in series) can be used to provide the high voltage power. The DC/DC converter and the power storage element BT1 are connected to each of the HV loads through 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 a combination of active and passive fuses (e.g., F1-F7) to protect against various failure 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>. For example, in some embodiments, fuses F2 EE1, F3 EE2, and F4 Xlink may correspond to fuses <NUM>, <NUM>, and <NUM> detailed in <FIG>.

Fuse F1 may be a pack fuse to disconnect the failing battery pack <NUM> from the rest of the HVPS system. In some embodiments, F1 may be a pyro-technical fuse. Upon failure of a battery pack <NUM>, the pyro-technical fuse F1 may receive a signal (e.g. from the associated battery management system) and blow the fuse F1. Therefore, further damage to other equipment (e.g. electric engines, EPUs, connected battery packs) can be avoided. Further, the electric engines <NUM> will still receive power from the paired battery packs <NUM> within the battery pack unit. For example, upon a battery pack failure, battery pack <NUM> pyro-technical fuse F1 may be blown, but electric engines <NUM>, <NUM>, <NUM>, and <NUM> may still receive power from battery pack <NUM>.

The arrangement of circuitry in the high voltage junction box (HVJB) <NUM> provides flexibility in charging by allowing for auxiliary loads and/or electric engines and actuators to be energized or de-energized in the charging process. For example, the battery pack <NUM> may be charged while the remaining HVPS circuitry remains disconnected. Charging contactors K4 positive and K4 negative may be closed to allow the battery pack <NUM> to charge. Meanwhile, main contactors K1 and K2 and pre-charge contactors (and/or relays) K3 and K5 may be open to prevent energizing the remaining HVPS circuitry. Further, the battery pack <NUM> may be charged while the auxiliary loads are connected but electric engines and actuators remain disconnected. Charging contactors K4 positive and K4 negative may be closed to allow the battery pack <NUM> to charge. Meanwhile, main contactors K1 and K2 may be closed after pre-charge contactor (and/or relays) K3 finishes pre-charging the auxiliary loads, and K4 EE may remain open. Further, the battery pack <NUM> may be charged while all loads are connected. Charging contactors K4 positive and K4 negative may be closed to allow the battery pack <NUM> to charge. Meanwhile, main contactors K1 and K2 may be closed after pre-charge contactors (and/or relays) K3 and K5 finish pre-charging the connected loads, and K4 EE may be closed.

In some embodiments, an input device may allow a person to select the charging mode of the aircraft. For example, a person 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 may receive a 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 a BMS <NUM> and the BMS <NUM> may control the contactors according to the requested charging mode.

<FIG> illustrates a diagram of a High Voltage Junction Box <NUM> (HVJB), consistent with embodiments of the present disclosure. In some embodiments, each battery pack <NUM> contains an HV distribution unit <NUM>, a Battery Management System (BMS <NUM>), and a Pyro-fuse Redundant Trigger board (PRT <NUM>) housed within the HVJB <NUM>. Each unit may be a hardware device, such as a computer, processor, or microprocessor. The BMS <NUM> may be configured to monitor voltages, temperatures, currents, and isolation resistances. The BMS <NUM> may control battery pack contactors and pyrotechnical fuses to protect against fault conditions. As further detailed below, the BMS <NUM> may communicate with various systems within and outside the HVJB <NUM>. The BMS <NUM> may include a Battery Management Unit (BMU <NUM>) which may receive voltage, current, resistance, and temperature sensing signals from the cell stack assembly <NUM> and/or the HV distribution unit <NUM>. The BMS <NUM> may further include Cell Management Units (CMUs) <NUM> to monitor the voltages of each set of <NUM> parallel cells (i.e., a <NUM>-7P cell group) connected in series in a <NUM>-7P cell block. The CMUs may also be used to monitor a <NUM>-7P cell block's temperature. The CMUs <NUM> obtain measurements for all the cell groups in the battery pack <NUM> and communicate the measurements to the BMU <NUM>.

The BMU <NUM> may monitor output current for each of the connected loads. The BMU <NUM> may be internally powered by the battery cell stack assembly <NUM> and continuously monitor the state of the battery even when it is not installed in the aircraft <NUM>. By monitoring the battery pack <NUM>, cell block, and cell group parameters, the BMU may protect against conditions that adversely affect safety or performance, such as overvoltage, undervoltage, overtemperature, under-temperature, loss of electrical isolation, short circuit, overcurrent, etc. The diagnostic function of the BMU <NUM> allows for fault detection and isolation through built-in-tests (BIT). In addition, the BMU <NUM> performs computation of the state of charge (SOC), state of health (SOH), failure condition (e.g. short circuit or overcurrent), state of power (SOP), state of energy (SOE) and state of temperature (SOT) of the battery pack <NUM>. The BMU <NUM> also controls and monitors bus pre-charging, provides fuse and contactor commands, and communicates with various systems within and outside the HVJB <NUM>.

HV distribution unit <NUM> in the HVJB <NUM> may contain HV contactors <NUM> and a combination of active and passive fuses (e.g., pyrotechnical fuses <NUM> and fuses <NUM>) to protect against overcurrent and short-circuit conditions. In some embodiments, the contactors <NUM> may correspond to one or more of switching devices K1-K7 (e.g. HV contactors) detailed in <FIG>. Similarly, the pyrotechnical fuses <NUM> and fuses <NUM>, may correspond to one or more fuses F1-F8 detailed in <FIG>. HV Distribution Unit <NUM> may further include (or receive information from) current sensors (e.g. resistor R3-R6, a hall effect sensor, shunt current sensor, or other sensor(s)).

In some embodiments, a pyro-fuse redundant trigger board (PRT <NUM>) may be located within HVJB <NUM>. While in other embodiments, BMS <NUM> may communicate with a PRT <NUM>, located outside the HVJB <NUM>. The BMS <NUM> may detect a failure event and send command signals to the PRT <NUM> for a corresponding pyro fuse driver to blow a fuse. For example, in some embodiments, HV Distribution Unit <NUM> may receive a sensor signal from a current sensor (e.g. resistor R3-R6) and provide information to the BMU <NUM> regarding the condition of the connected loads (e.g. a voltage, current, or temperature) at a point in the HVPS system. Based on the received information, the BMU <NUM> may determine a failure condition (e.g. because the value is outside a predetermined range) and send a command to PRT <NUM> to blow an associated pyrotechnical fuse. Therefore, the fault condition can be disconnected from the rest of the HVPS circuitry, protecting the remaining devices and wiring. In some embodiments the BMU <NUM> may directly monitor the sensors instead of receiving information through HV Distribution Unit <NUM>.

In some embodiments, battery packs <NUM> may be in communication with each other, e.g. through BMS <NUM>. The battery packs <NUM> may use information regarding the state of one or more paired battery packs <NUM> in a battery pack unit to help determine whether an overcurrent condition has occurred. For example, a battery pack <NUM> may determine an expected operation range (e.g. voltage, current etc.) based on the state of the battery pack and the communicated state of battery packs <NUM> within the battery pack unit. In some embodiments, HVJB <NUM> may further provide a redundant active trigger board configured to enable the pyro fuse driver to activate one or more pyrotechnical fuses when the BMS <NUM> fails to enable the pyro fuse driver. See <CIT> incorporated by reference.

The Control MCU (CCU <NUM>) in the charge port assembly <NUM> may interface with the external battery charger and communicate with the BMUs <NUM> on the six installed battery packs <NUM>. This unit may be a hardware device, such as a computer, processor, or microprocessor. In some embodiments, the CCU <NUM> may be a single PCBA with one microcontroller that manages overall power delivery to each battery pack <NUM> when charging. As shown in <FIG>, the CCU <NUM> may perform the handshake between the Ground Charging Subsystem <NUM> and the BMUs <NUM> and may command the BMUs <NUM> to open or close contactors <NUM>, such as contactors K6-K7 detailed in <FIG>. The CCU <NUM> may perform active detection and protection features for overvoltage protection. The BMUs <NUM> in each battery pack <NUM> may retain full control and continuously monitor their battery packs <NUM> during charging operations.

<FIG> illustrates a diagram of a Charge Port Assembly (CPA <NUM>), consistent with embodiments of the present disclosure. Charge Port Assembly <NUM> includes a Charge Port <NUM>, providing for communication connection through power line communication <NUM> and HV power transfer through HV power channel <NUM>. In some embodiments, Charge Port <NUM> may be a JI <NUM> Type <NUM> charge port including various pins and connection points to allow for connection to a Ground Service System (GSS <NUM>) (e.g. through a plug). The Charge Port <NUM> may include one or more proximity pins to detect a high voltage connection between the GSS <NUM> and the Charge Port <NUM>. Upon detecting a connection with the GSS <NUM>, the Charge Port <NUM> may engage a latch that prevents the high voltage power <NUM> from being disconnected under a charged load. Following completion of the charging, the Charge Port <NUM> may automatically unlatch the connection or enable manual unlatching.

The Charge Port Assembly <NUM> may include a Charge Control Unit (CCU <NUM>) in communication with the Charge Port <NUM>, e.g. through communication line <NUM>. The CCU <NUM> may further provide latch control <NUM>, illumination changes <NUM>, and to monitor and respond to a temperature <NUM> of various components. The CCU <NUM> may monitor a temperature on an inlet side of the charge port <NUM>. If the temperature gets too high, then the CCU <NUM> may command the Ground Service System <NUM> to abort the charge. The CCU <NUM> receives status updates from battery packs <NUM> and provides commands to battery packs <NUM> to control their charge level by opening and closing battery pack charge contactors (e.g. K6-K7 in <FIG>). As detailed with reference to <FIG> above, in some embodiments, the CCU <NUM> may communicate with each battery pack's Battery Management System (BMS <NUM>), e.g. through a Battery Management Unit (BMU <NUM>). The BMS <NUM> may send battery pack information to the CCU <NUM>, including information on a state of battery pack connection (e.g. whether the battery pack is connected to the HVPS system), state of charge (SOC), state of health (SOH), failure condition (e.g. short circuit or overcurrent), state of power (SOP), state of energy (SOE), and state of temperature (SOT).

The CCU <NUM> may provide commands to the BMS <NUM> to open or close battery pack charge contactors. In some embodiments, each battery pack <NUM> may have a separate low voltage CAN communication line connecting the battery pack <NUM> to the CCU <NUM>. In some embodiments, a CAN communication line may be shared between one or more battery packs <NUM> in a battery pack unit. For example, HV battery packs <NUM> and <NUM> may communicate with CCU <NUM> through CAN <NUM>. HV battery packs <NUM> and <NUM> may communicate with CCU <NUM> through CAN <NUM>. HV battery packs <NUM> and <NUM> may communicate with CCU <NUM> through CAN <NUM>. As further detailed below, CCU <NUM> may make various power supply and cooling requests of the GSS <NUM> (e.g. through charge port <NUM>) based on the information received from the battery packs <NUM>.

Charge Control Unit <NUM> may determine battery pack charge contactor commands based on a variety of criteria. In some embodiments, CCU <NUM> may determine the required battery pack charge levels based on flight information. For example, in some embodiments CCU <NUM> may receive flight information from GSS <NUM>, e.g. through communication lines <NUM> and <NUM>. GSS <NUM> 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, Charge Port Assembly <NUM> may provide for a direct wired or wireless connection to a computer, laptop, ipad, mobile device to directly receive flight information. In some embodiments, CCU <NUM> may receive flight information from the aircraft's flight control system <NUM>.

Flight information may include flight mission information, such as a location of the destination, a distance to the next destination, or an expected flight time required to get to the next destination. Flight mission information may include a type of flight expected. For example, flight mission information may include a duration or distance to be covered in each flight mode. In some embodiments, flight modes may include winged-flight, thrust and lift assisted flight, thrust assisted flight, and lift assisted flight. In some embodiments, flight mission information may include an expected EPU output throughout the flight, e.g. as a unit of power or percentage of max EPU power. In some embodiments, flight mission information may be provided for each EPU on an aircraft.

Flight mission information may include information on predicted weather conditions throughout the flight. Weather conditions may include temperatures, pressures, wind conditions, and precipitation expected throughout the flight. Flight mission information may include an expected weight of an aircraft, e.g. based on the number of passengers or an amount of cargo. The weight of an aircraft may be predicted or measured (e.g. if the aircraft is charging with passengers or cargo on board).

Flight information may include historical battery information. For example, in some embodiments, battery information may include historical battery consumption of each battery pack on a particular flight path. The battery information may further include details on flight modes, weight, and weather, for the Charge Control Unit <NUM> to determine its relevance to the flight mission ahead.

Further, flight information may be received and analyzed for multiple subsequent flights. In some embodiments, if an aircraft will take multiple trips without the ability to re-charge, flight information may be gathered and analyzed for all subsequent flights to ensure the aircraft has sufficient charge for each trip. In some embodiments, an aircraft may have time to partially re-charge before a subsequent trip. Therefore, flight information may include information on the subsequent trip and information on the amount of re-charging that is available between trips. By receiving this information, the CCU <NUM> may ensure that the battery packs <NUM> have enough charge to support a sufficient portion of the subsequent trip. The CCU <NUM> may use the flight information to determine a required charge level required of each battery pack <NUM>.

Charge Control Unit (CCU <NUM>) may determine battery pack charge contactor commands based on the current state of each battery pack <NUM> received from the BMS <NUM>, including a state of energy and/or state of charge of each battery pack <NUM>. The CCU <NUM> may determine how much additional charge is necessary to meet the required charge level based on each battery pack's current charge level. Further, in some embodiments, the CCU <NUM> may consider the battery pack configuration when charging the battery packs <NUM>. The CCU <NUM> may determine to charge each battery pack <NUM> within a battery pack unit to the same charge level. Therefore, the CCU <NUM> may charge all battery packs <NUM> in a battery pack unit to the highest charge level required of any battery packs <NUM> within the unit. As the battery pack <NUM> charges, the CCU <NUM> may continue to receive updates on each battery pack's charge level and keep the battery pack charge contactors closed to enable charging until the required charge level is reached.

Further, Charge Control Unit (CCU <NUM>) may determine battery pack charge contactor commands based on a failure condition, state of health, or state of temperature received from the BMS <NUM>. In some embodiments, CCU <NUM> may open a contactor to a battery pack <NUM> (disabling charging) based on receiving information that a battery pack <NUM> has failed (e.g. experienced a short circuit or overcurrent condition). Further, the CCU <NUM> may open a contactor to a battery pack <NUM> (disabling charging) based on the battery pack state of health dropping below a set level or based on the battery pack temperature exceeding a set level. The CCU <NUM> may continue to monitor failure condition, state of health, or state of temperature from the battery pack <NUM>, and close the contactor (enabling charging) when the conditions are remedied.

Charge Control Unit <NUM> may send cooling commands to the GSS <NUM>, e.g. through communication lines <NUM> and <NUM>, based on the state of temperature information received from the battery packs <NUM>. In some embodiments, the CCU <NUM> may send a required battery pack temperature or a required coolant flow rate. The Ground Charging Subsystem <NUM> may communicate this information with a Thermal Conditioning Subsystem <NUM>. The Thermal Conditioning Subsystem <NUM> may control one or more condensers and associated coolant control valves to achieve the cooling requirements.

The Charge Control Unit (CCU <NUM>) may signal the state of the battery packs <NUM> throughout the charging process to a charging attendant. In some embodiments, the CCU <NUM> may signal a problem (e.g. a battery pack failure, poor health, or excess temperature) through the illumination line <NUM>. For example, in some embodiments, a light may be turned on or change colors to indicate the problem. Alternatively, or additionally, the CCU <NUM> may communicate the details of the problem (e.g. type of problem, relevant battery pack(s) etc.) to the Ground Service System <NUM> through communication lines <NUM> and <NUM>. Ground Service System <NUM> may provide these details through a display, computer, laptop, ipad, mobile device, or any other device capable capable of communicating the information to a charging attendant.

Charge Control Unit (CCU <NUM>) may determine that each battery pack <NUM> has reached the required charge level and signal charge completion to the Ground Service System <NUM>. Upon determining that no charge is being received from the GSS <NUM>, the CCU <NUM> may provide a signal to the charge port, e.g. through latch control <NUM>, to automatically unlatch the connection to the GSS <NUM> or to allow for manual unlatching of the connection.

<FIG> illustrates a flow chart for detecting an emergency responder, consistent with embodiments of the present disclosure. In some embodiments, this process is performed by each battery management system <NUM> of the battery packs <NUM>. At step <NUM>, a processor, receives acceleration information. In some embodiments, acceleration information may be received directly from sensors (e.g. an accelerometer), while in other embodiments acceleration information may be received from a different processor, such as one associated with a flight control system of the aircraft. At step <NUM>, the processor, receives a High Voltage Interlock Loop (HVIL) continuity status (e.g. from a Battery Management System (BMS <NUM>)) indicating whether or not a low voltage emergency cut loop has been cut. For example, a BMS <NUM> determines that a cut loop has been cut based on detecting a loss of current. The information gathered in steps <NUM> and <NUM> may be received sequentially or simultaneously. Further, in some embodiments the information gathered may include a time stamp indicating when it was collected. While in other embodiments, the processor may assign a time based on when it received the information.

At step <NUM>, the processor may determine whether an emergency responder performed a cut of the low voltage emergency cut loop. The processor may make this determination based on the acceleration information and the 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, then it is determined that an emergency responder cut the low voltage emergency cut loop cut loop. However, if an HVIL continuity status indicates a cut loop at an earlier time than the acceleration information indicates a crash, then an emergency response is not detected. Further if either the acceleration information doesn't indicate a crash or the HVIL continuity status does not indicate a cut, then an emergency response is not detected. At step <NUM>, if it is determined that an emergency responder performed the cut then the processor sends 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 <NUM> to blow based on which battery pack <NUM> is associated with the cut loop. For example, in some embodiments a cut loop <NUM> may be connected to a battery pack <NUM>. The processor may determine an emergency responder cut loop <NUM> and the processor may instruct battery pack <NUM> to blow the battery pack <NUM> pyrotechnical fuse, such as fuse F1 in <FIG>. In some embodiments, based on determining an emergency responder cut any of the loops, the processor may blow the pyrotechnical fuse associated with the battery pack <NUM> and any connected battery pack <NUM>. For example, referencing <FIG>, based on determining that an emergency responder cut loop <NUM> associated with battery pack <NUM>, the processor may blow the pyrotechnical fuses associated with battery packs <NUM> and <NUM>. In some embodiments, based on determining an emergency responder cut any of the loops, the processor may blow the pyrotechnical fuse associated with all the battery packs <NUM>.

At step <NUM>, the processor may determine whether the crash detection was false. The processor may determine that acceleration information indicates a crash, but the HVIL continuity status indicates that there is no cut loop. Further, the processor may gather, or have available, information on whether the flight control system is in ground mode. If the processor determines that the aircraft flight control system is in ground mode, the processor may determine that the crash detection was false. However, if the processor determines that the flight control system is not in ground mode (e.g. in fly mode), then a false crash will not be 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.

At step <NUM>, if it is determined that the crash detection was false, Condition <NUM> will be reset to indicate no crash detected and the processor will re-gather acceleration information. At step <NUM>, if it is not determined that the crash detection was false, condition <NUM> will not be reset and the processor will continue to monitor whether the HVIL continuity status indicates a cut loop at Step <NUM> condition <NUM>.

<FIG> illustrates a plan view diagram for routing cut loop wiring through the tail of an eVTOL aircraft, consistent with embodiments of the present disclosure. As detailed above, each cut loop may be connected to a single battery pack <NUM>. Therefore, <NUM> cut loops may be be routed from the battery packs <NUM> located in the wings, or elsewhere, to the tail of the plane. This routing ensures that the cut loops are accessible to be cut in the tail of the aircraft away from the high voltage power system running between the batteries, electric engines, and other aircraft devices towards the front of the aircraft. A first responder can cut one or more loops to de-energize the battery packs <NUM> without risking cutting into an energize high voltage line, thereby increasing safety. In some embodiments, each cut loop may be routed separately. In some embodiments, the cut loops may be routed with one or more battery packs <NUM> (e.g. in a bundle). For example, cut loops associated with connected battery packs may be bundled together or cut loops associated with a wing of the plane may be bundled together. In some embodiments, the cut loops for the battery packs <NUM> may all be routed together in a single bundle.

Claim 1:
A method for aircraft battery management, the method comprising:
receiving, by a battery management system (<NUM>), aircraft movement information of an aircraft;
the method being characterized by further comprising:
detecting, by the battery management system (<NUM>), that the movement information indicates a potential crash;
detecting, by the battery management system (<NUM>), a loss of current in at least one low voltage wire; and
blowing, by the battery management system (<NUM>), a battery pack fuse (<NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>) of at least one battery pack (<NUM>) configured to supply high voltage power to disconnect supply of the high voltage power by the at least one battery pack.