High voltage battery architecture

Abstract

A method of controlling aircraft power distribution comprises: receiving aircraft state information of an aircraft, detecting whether the aircraft state information indicates a crash, and disconnecting supply of high voltage power to the aircraft by at least one battery upon detecting both that the aircraft state information indicates a crash and a loss of continuity in at least one low voltage wire.

Description

HIGH VOLTAGE BATTERY ARCHITECTURE CROSS-REFERENCE TO RELATED APPLICATIONS[1] This application claims priority to PCT Application PCT / US23 / 79690 (Attorney Docket No. 16163.0018-00304) filed November 14, 2023, titled “High Voltage Battery Architecture” which in turn claims priority to and the benefit of U.S. Provisional Application No. 63 / 383,660, filed November 14, 2022, titled “Systems and Methods for Improved Battery Assemblies for eVTOL Aircraft” (Attorney Docket No. 16163.6005-00000). The entire contents of the aforementioned applications are incorporated by reference herein for all purposes.TECHNICAL FIELD[2] This disclosure relates generally to the field of powered aerial vehicles. More particularly, and without limitation, the present disclosure relates to innovations in tilt-rotor aircraft that use electrical propulsion systems. Certain aspects of the present disclosure generally relate to configuration and control of high voltage power systems for aircrafts.BACKGROUND[3] The inventors here have recognized several problems associated with powering an electric aircraft, including a tilt-rotor aircraft that uses electric or hybrid-electric propulsion systems (hereinafter referred to as electric propulsion units or “EPUs”). 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 forward thrust support. Therefore, there it is critical to provide for redundancy in the aircraft’s high voltage power system to avoid a single point of failure. It is also critical to ensure a fault or failure condition does not propagate and damage other critical aircraft components. Additionally, in the event of a crash, there is a need to allow a first responder to shut off the high voltage power system quickly and safely (e.g., to avoid shock hazards and / or to quickly stop propeller rotation).SUMMARY[4] The present disclosure generally relates to an electrical system for an aircraft. More particularly, and without limitation, the present disclosure relates to innovations in an aircraft with electric propulsion units. Certain aspects of this disclosure relate to an electrical system that provides for both redundancy and fault tolerance. Other aspects of the present disclosure relate to an electrical system that allows a first responder to shut off high voltage powerquickly and safely using low voltage cut loops. Further aspects of the present disclosure relate to routing the low voltage cut loops in a manner that avoids a single point of failure affecting multiple systems.[5] One aspect of the present disclosure is directed to a method of controlling aircraft power distribution, comprising: receiving aircraft state information of an aircraft, detecting whether the aircraft state information indicates a crash, and disconnecting supply of high voltage power to the aircraft by at least one battery upon detecting both that the aircraft state information indicates a crash and a loss of continuity in at least one low voltage wire.[6] Another aspect of the present disclosure is directed to an electrical system for an aircraft, comprising: a first battery, a second battery, a first low voltage cut loop wire connected to the first battery and a second low voltage cut loop wire connected to the second battery, wherein the first low voltage cut loop is spaced apart from the second low voltage cut loop.[7] Another aspect of the present disclosure is directed to an electrical system for an aircraft, comprising: a first paired battery pack unit comprising a first battery and a second battery, the first battery electrically connected to the second battery via a first high voltage bus and a second paired battery pack unit comprising a third battery and a fourth battery, the third battery electrically connected to the fourth battery via a second high voltage bus, wherein the first high voltage bus and second high voltage bus are electrically separate from one another. The electrical system further comprising a first low voltage cut loop connected to the first battery and a second low voltage cut loop connected to the third battery.[8] Another aspect of the present disclosure is directed to an aircraft comprising a wing, a fuselage, a cabin area section of the fuselage, high voltage wiring, battery packs installed in the wings, wherein walls and floor of the cabin area are free of high voltage wiring and wherein the wing includes high voltage wiring.BRIEF DESCRIPTIONS OF FIGURES[9] Figure 1A illustrates an example electric aircraft, consistent with embodiments of the present disclosure.

[0010] Figure IB illustrates another example electric aircraft, consistent with embodiments of the present disclosure.

[0011] Figure 1C illustrates an electric engine 110 with two partial motors, consistent with embodiments of the present disclosure.

[0012] Figure ID illustrates a diagram of a high voltage power distribution system for an electric aircraft, consistent with embodiments of the present disclosure.

[0013] Figures IE, IF, 1G, 1H, II and 1J illustrate exemplary top plan views of aircraft, consistent with disclosed embodiments.

[0014] Figure 2 illustrates a circuit diagram for a High Voltage Junction Box (HVJB), consistent with embodiments of the present disclosure.

[0015] Figure 3A illustrates a diagram of a High Voltage Junction Box (HVJB), including a battery management system, consistent with embodiments of the present disclosure.

[0016] Figure 3B illustrates a central battery management system, consistent with embodiments of the present disclosure.

[0017] Figure 4A illustrates a flow chart for detecting an emergency response, consistent with embodiments of the present disclosure.

[0018] Figure 4B illustrates another flow chart for detecting an emergency response, consistent with embodiments of the present disclosure.

[0019] Figures 5A and 5B illustrate areas where the high voltage wiring is located on the aircraft, consistent with embodiments of the present disclosure.

[0020] Figure 5C illustrates high voltage wiring routed through the wing of an aircraft, consistent with embodiments of the present disclosure.

[0021] Figures 6A and 6B illustrate plan view diagrams for routing low voltage emergency cut loop wiring through the tail of an aircraft, consistent with embodiments of the present disclosure.

[0022] Figure 6C illustrates a profile view diagram for routing cut loop wiring through the tail of an aircraft, consistent with embodiments of the present disclosure.DETAILED DESCRIPTION

[0023] The present disclosure addresses components of an electric aircraft. For example, the aircraft may be an electric vertical takeoff and landing (eVTOL) aircraft intended for frequent (e.g., over 50 flights per workday), short-duration flights (e.g., less than 100 miles per flight) over, into, and out of densely populated regions. Accordingly, it is important that the aircraft has a variety of safety and backup mechanisms. For example, a distributed propulsion system can avoid the risk of a single point of failure and enhance safety.

[0024] Disclosed embodiments provide new and improved configurations of aircraft components that are not observed in conventional aircraft, and / or identify design criteria for components that differ from those of conventional aircraft. Such alternate configurations and design criteria, in combination with addressing drawbacks and challenges with conventional components, yielded the embodiments disclosed herein for various configurations and designs of electric aircraft components.

[0025] In some embodiments, the electric aircraft (e.g., an 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 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.

[0026] 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.

[0027] It is critical that an electric aircraft’s power system includes redundancy, fault tolerance, and emergency shut off capabilities. Some of the disclosed embodiments achieve these goals, and others, 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.

[0028] Additionally, some of the disclosed embodiments provide an ability for a first responder to shut off the high voltage power system quickly and safely based on detecting an aircraft state indicates a crash and a first responder has cut a low voltage cut loop. Further, the cut loop(s) may be routed through the tail of the aircraft to provide separation from high voltage circuits and increase safety for the first responder. The cut loops may be spaced in a manner that prevents a single failure point from impacting multiple battery packs and associated EPUs (e.g., electric engines).

[0029] Figure 1A illustrates an example electric aircraft (e.g., an eVTOL aircraft), consistent with embodiments of the present disclosure. As shown in Fig. 1 A, in some embodiments, the distributed electrical propulsion system of the eVTOL aircraft 100 may include twelve electric engines 110, which may be mounted on booms forward and aft of the main wings of aircraft 100. Forward electric engines 110 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 thrust). Forward electric engines 110 may be of a clockwise type orcounterclockwise type in terms of direction of propeller rotation. Aft electric engines 110 may be fixed in a vertically oriented position (e.g., to generate vertical thrust), and may also be of a clockwise type or counterclockwise type in terms of direction of propeller rotation.

[0030] Aircraft 100 may possess various combinations of forward and aft electric engines 110. For example, in some embodiments, aircraft 100 may possess six forward electric engines 110 and six aft electric engines 110. In some other embodiments, aircraft 100 may include four forward electric engines 110 and four aft electric engines 110, or any other combination of forward and aft engines 110. In some other embodiments, the number of forward electric engines and aft electric engines are not equivalent.

[0031] In some embodiments, for a vertical takeoff and landing (VTOL) mission, forward electric engines 110 as well as aft electric engines 110 may provide vertical thrust during takeoff and landing. During flight phases where aircraft 100 is in forward flight, forward electric engines 110 may provide forward thrust (e.g., in a horizontal direction), while the propellers of the aft electric engines 110 may be stowed at a fixed position in order to minimize drag. Aft electric engines 110 may be actively stowed with position monitoring.

[0032] In some embodiments, in a conventional takeoff and landing (CTOL) mission, forward electric engines 110 may provide forward thrust for wing-borne take-off, cruise, and landing. In some embodiments, aft electric engines 110 may not be used for generating thrust during a CTOL mission and the aft propellers may be stowed in place.

[0033] 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 phase to a mostly horizontal direction during forward-flight phase. 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.

[0034] As shown in Fig. 1 A, aircraft 100 may be configured with a distributed electric propulsion system enabling vertical flight, forward flight, and transition. The forward 6 electric engines 110 (which are numbered 1-6 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 6 electric engines 110 (which are numbered 7-12 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. Aircraft 100 will be equipped with advanced cockpit avionics, aflight management system, and the sensors necessary to support the intended operations and system functions.

[0035] 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 110, which in turn converts HV power into mechanical shaft power which is used to rotate a propeller. As mentioned above, an aircraft 100 as described herein may possess multiple electric engines 110 which are boom-mounted forward and aft of the wing. The amount of thrust each electric engine 110 generates may be governed by a torque command from the Flight Control System (FCS) over a digital communication interface to each electric engine 110.

[0036] In some embodiments, aircraft 100 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 aircraft 100 and configured to distribute the stored electrical energy to other systems on aircraft 100, including the electrical propulsion system (EPS) for converting electrical power into mechanical rotational shaft power to generate thrust. As shown in Fig.1 A, the HVPS system of aircraft 100 may include six battery packs 120 (which are numbered 1-6 from left to right) installed within the battery bays in the wing of aircraft 100. In some embodiments, six battery packs 120 may have the identical design, to simplify the design, manufacturing, and logistics. Battery packs 120 may power one or more electric engines 110. While six battery packs 120 are shown, aircraft 100 may have any number of battery packs 120.

[0037] In some embodiments, a single battery pack 120 may be electrically connected to, and power, multiple electric engines 110. For example, in some embodiments, a battery pack 120 may power an electric engine 110 on either side of a longitudinal axis running through the center of the aircraft. In some embodiments a battery pack 120 may power an electric engine 110 on either side of a horizontal axis running through the wing. In some embodiments, as shown in Fig. 1 A, a battery pack 120 may power two diagonally opposing 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. Therefore, upon a loss of a battery pack 120, the impact to roll or pitch moments can be reduced because the loss of lift is balanced. In some embodiments, battery packs 120 may power different arrangements of electric engines 110 to reduce roll, pitch, or yaw moments that may be caused by a loss ofbattery pack 120. For example, in some embodiments, battery packs 120 may be connected to electric engines 110 in any manner that balances lift and / or thrust across the longitudinal and horizontal axis of the aircraft.

[0038] Further, the HVPS system includes a cross-link 130 possessing at least one fuse allowing for pairing of two or more battery packs 120. Through the cross-link, power for the electric engines 110 can be shared among the paired battery packs 120. Therefore, multiple battery packs 120 can simultaneously power multiple electric engines 110. This arrangement provides for redundancy and avoids a single point of failure because each paired battery 120 may act as a backup for the other(s). Upon failure of a battery pack 120, one or more connected battery packs 120 may continue powering the failed battery pack’s connected electric engines 110.

[0039] In some embodiments, as shown in Fig. 1 A, a pair of battery packs 120 may include two battery packs 120. In some embodiments, a pair of two battery backs 120 may power a total of four electric engines 110. For example, battery pack 1, providing power to electric engines 1 and 12, may be cross-linked to battery pack 4, providing power to electric engines 4 and 9. Battery pack 2, providing power to electric engines 5 and 8, may be cross linked to battery pack 5, providing power to electric engines 2 and 11. Battery pack 3, providing power to electric engines 3 and 10, may be cross linked to battery pack 6, providing power to electric engines 6 and 7.

[0040] Figure IB illustrates another example electric aircraft (e.g., an eVTOL aircraft), consistent with embodiments of the present disclosure. In some embodiments, electric engines 110 may include multiple motor stages that are each independently powered by different battery packs 120 so that should one battery pack 120 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 110 may include two partial motors. For example, battery pack 1 may power first partial motors on electric engines 1, 6, 7, and 12. Battery pack 6 may power second partial motors on electric engines 1, 6, 7, and 12. In some embodiments, different configurations may be used. For example, battery pack 1 may power first partial motors on electric engines 1, 4, 9, and 12. Battery pack 6 may power second partial motors on electric engines 1, 4, 9, and 12.

[0041] Figure 1C illustrates an electric engine 110 with two partial motors 191a and 191b, consistent with embodiments of the present disclosure. Partial motors 191a and 191b may be powered by different battery packs 120, as shown in Fig. IB above. Two partial motors 191a and 191b can operate independently to drive blades of an EPU and can operatesimultaneously to drive the blades at a higher power. 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 110 has an electrically separate backup.

[0042] Alternate Aircraft and Battery Pack Arrangements

[0043] The above aircraft and battery pack configurations are provided as examples, but the aircraft may include a different configuration of electric propulsion units (e.g., as shown below with reference to Fig. IE-1 J), battery packs, battery pack connections, and battery pack cross link combinations. In some embodiments, each battery pack may power an individual electric propulsion unit (e.g., electric engine). For example, an aircraft may have four, six, eight, ten, twelve, or any number of electric propulsion units and the number of battery packs may match the number of electric propulsion units. In some embodiments, each battery pack may power only one electric propulsion unit and may be electrically separate from all other battery packs. In some embodiments, each battery pack may power one or more partial motors and each electric propulsion unit (e.g., electric engine) may include two or more partial motors. Therefore, each electric propulsion unit may have a backup power source, but the battery packs may still be electrically separate.

[0044] In some embodiments, each battery pack may power multiple electric propulsion units (e.g., electric engines). As described above, battery packs may power sets of electric engines that are symmetrical across one or more axes of symmetry. In some embodiments, a battery pack may power electric propulsion units (e.g., electric engines) that are symmetrical across an aircraft’s longitudinal axis, lateral axis, or both. For example, as described above, in some embodiments, different battery packs may power diagonally symmetric electric propulsion units (e.g., electric engines).

[0045] In some embodiments, a battery pack may power more than two electric propulsion units (e.g., electric engines). In some embodiments, a battery pack 120 may power two or more sets of diagonally symmetric electric propulsion units. In some embodiments, the set of electric engines powered by a battery pack may include an inboard diagonally symmetric pair of electric propulsion units and an outboard diagonally symmetric pair of electric propulsion units (e.g., electric engines). In some embodiments, a battery pack may power four or more electric propulsion units (e.g., electric engines) in a configuration that is symmetrical across the longitudinal axis of symmetry.

[0046] In some embodiments, some or all of the battery packs are interconnected. As described above, a cross-link may allow each battery pack to act as backup power foranother. For example, in some embodiments, a first battery pack may directly power a first number electric propulsion units (e.g., electric engines) and a second battery pack may directly power a second number of electric propulsion units (e.g., electric engines). The first and second battery packs 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 propulsion units (e.g., electric engines) 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.

[0047] In some embodiments, a battery pack unit may comprise three battery packs, wherein each battery pack powers a number of electric propulsion units (e.g., electric engines). For example, in some embodiments, each battery pack may power two diagonally symmetric electric engines. Therefore, each battery pack unit may power a total of six electric engines and each electric engine has two battery pack backups. In some embodiments, each battery pack in the battery pack unit may power four electric engines, comprising two sets of diagonally symmetric electric engines. Therefore, each battery pack unit may power a total of twelve electric engines and each electric engine has two battery pack backups.

[0048] In some embodiments, a battery pack unit may comprise four battery packs, wherein each battery pack powers a number of electric propulsion units (e.g., electric engines). For example, in some embodiments, each battery pack may power two diagonally symmetric electric engines. Therefore, each battery pack unit may power a total of eight electric engines and each electric engine has three battery pack backups. In other embodiments, each battery pack in the battery pack unit may power four electric engines, comprising two sets of diagonally symmetric electric engines. Therefore, each battery pack unit may power sixteen electric engines and each electric engine has three battery pack backups.

[0049] In some embodiments, all battery packs are connected to a common bus. In some embodiments, the common bus may form a circular power supply, providing for additional redundancy in connections, while in other embodiments the common bus may not form a circular power supply.

[0050] In some embodiments, electric engines comprise a single motor that is powered by the one or more battery packs. In some embodiments, each electric engine may include two or more partial motors and the battery packs may power partial motors. In some embodiments, any of the electric engine powering configurations described above may include powering partial motor(s) of battery pack(s).

[0051] Different configurations of battery packs, electric propulsion units (e.g., electric engines), battery pack to electric propulsion unit connections, and battery pack cross link combinations may be chosen to best balance aircraft power needs, system redundancy, and fault tolerance. As described below, in some embodiments, the low voltage cut loop routing, bundling, and / or spacing is configured based on the arrangement of battery packs and / or the battery pack cross link combinations. Further, as described below, battery pack disconnect(s) following an emergency response may be based on the arrangement of battery packs and / or the battery pack cross link combinations.

[0052] Figure ID illustrates a diagram of a high voltage power system for an electric aircraft (e.g., an eVTOL), consistent with embodiments of the present disclosure. As shown, an electric aircraft may include a battery assembly comprising electrically separate battery pack units (e.g. 160, 162, and 164). Each battery pack unit may include battery packs 120 that are 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 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 120 that power electric engines 110 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., <±5%, <±10%, <±15%, <±20%, or <±25% asymmetry) with respect to roll, pitch, and / or yaw of the aircraft. In some embodiments, the battery pack units may include battery packs 120 to reduce an overall amount of high voltage wiring between the battery packs. In some embodiments, the battery pack units may include battery packs 120 to minimize power requirements.

[0053] In some embodiments, as shown in Fig. ID, the HVPS system may comprise 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. Therefore, each battery pack unit may include two paired battery packs 120 that simultaneously power four electric engines 110. Upon the failure of one battery pack 120 in a battery pack unit, the other paired battery pack 120 will continue powering the four electric engines.

[0054] In some embodiments, each battery pack units 160, 162, 164 may include a high voltage bus to cross-link battery packs 120 within the battery pack unit. In some embodiments, cross-link 130 connects two high voltage channels, each feeding one or more electric engines 110. For example, in some embodiments, cross link 130 may be connected to each battery pack’s high voltage channel before the channel splits to power multiple electric engines 110 (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).

[0055] In some embodiments, each cross link 130 may include at least one fuse to disconnect the cross-link upon a failure of the cross-link. For example, fuses 131, 132, and 134 may be located on the cross-link connection of the positive high voltage channels in battery pack units 160, 162, and 164. In some embodiments, the fuses may be pyro-technical fuses. As further detailed below, a battery management system of a connected battery pack 120 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, electric engines 110 will still receive power from the paired battery pack 120 in the battery pack unit. For example, upon a cross-link failure, pyro-technical fuse 131 may be blown, but electric engines 1 and 12 will still receive power from battery pack 1, and electric engines 4 and 9 will still receive power from battery pack 4. As further described below, in some embodiments, one or more cross-link fuses may be blown upon detection of an emergency response.

[0056] In some embodiments, there may be additional pyrotechnical fuses on the cross-link connection of the negative high voltage channels. For example, pyrotechnical fuses 170, 172, and 174 may be located on the cross-links in battery pack units 160, 162, and 164, 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.

[0057] 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 109, 111, 112, and 113 may be located on the high voltage channel 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.

[0058] 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 110 in the battery pack unit will still receive power from the connected battery packs 120. For example, upon a failure in a device or wiring downstream of pyrotechnical fuse 109, the pyrotechnical fuse 109 may be blown, but electric engines 12, 4, and 9 will still receive power from battery packs 1 and 4. 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. As further described below, in some embodiments, one or more load disconnection devices may be blown upon detection of an emergency response.

[0059] In some embodiments, the HVPS system may include a high voltage charging channel allowing all battery packs 120 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 140, 142, 144, 146, 148, and 150 may provide for disconnection of battery packs 1, 4, 5, 2, 3, and 6, 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 141, 143, 145, 147, 149, and 151 may provide for disconnection of battery packs 1, 4, 5, 2, 3, and 6, respectively.

[0060] Figs. IE, IF, 1G, 1H, Hand 1J illustrate exemplary top plan views of aircraft, consistent with disclosed embodiments.

[0061] Figs. 1E-1J are illustrations of a top plan view of exemplary aircrafts, consistent with embodiments of the present disclosure. There may be a number of design considerations (cost, weight, size, performance capability etc.) that may influence the number and / or combination of tilt and lift propellers in an aircraft. The number and orientation of propellers may affect the number of battery packs and connections between battery packs (e.g., to achieve controllability and / or stability upon electric failure).

[0062] Fig. IE illustrates an arrangement of electric propulsion units, consistent with embodiments of the present disclosure. Referring to Fig. IE, the aircraft shown in the figure may be a top plan view of an exemplary aircraft. The aircraft may include twelve electric propulsion systems distributed across the aircraft. In some embodiments, a distribution of electric propulsion systems may include six forward electric propulsion systems (165, 166, 167, 168, 169, and 170) and six aft electric propulsion systems (171, 172, 173, 174, 175, and, 176). In some embodiments, the six forward electric propulsion systems may be operatively connected to tilt propellers and the six aft electric propulsion systems may be operatively connected to lift propellers. In other embodiments, the six forward electric propulsion systems and a number of aft electric propulsion systems may be operatively connected to tilt propellers and the remaining aft electric propulsion systems may be operatively connected to lift propellers. In other embodiments, all forward and aft electric propulsion systems may be operatively coupled to tilt propellers.

[0063] Fig. IF illustrates an alternate arrangement of electric propulsion units, consistent with embodiments of the present disclosure. Referring to Fig. IF, the aircraft shown in the figure may be a top plan view of an exemplary aircraft. The aircraft may include eight electric propulsion systems distributed across the aircraft. In some embodiments, a distribution of electric propulsion systems may include four forward electric propulsion systems (177, 178, 179, and 180) and four aft electric propulsion systems (181, 182, 183, and 184). In some embodiments, the four forward electric propulsion systems may be operatively connected to tilt propellers and the four aft electric propulsion systems may be operatively connected to lift propellers. In other embodiments, the four forward electric propulsion systems and a number of aft electric propulsion systems may be operatively connected to tilt propellers and the remaining aft electric propulsion systems may be operatively connected to lift propellers. In other embodiments, all forward and aft electric propulsion systems may be operatively coupled to tilt propellers.

[0064] Fig. 1G illustrates an alternate arrangement of electric propulsion units, consistent with the embodiments of the present disclosure. Referring to Fig. 1G, the aircraft may be a top plan view of an exemplary aircraft. In some embodiments, the aircraft may include ducted fans operably connected to the electric propulsion systems. In some embodiments the aircraft may include a bank of ducted fans on each wing of the aircraft and the bank of ducted fans may be connected to tilt together (e.g., between lift and forward thrust configuration). In some embodiments the aircraft includes a left and right front wing and a left and right rear wing. In some embodiments, each wing of the aircraft includes a bank of connected ducted fans. In some embodiments, each bank of connected ducted fans are tiltable (e.g., between lift and forward thrust), while in other embodiments only the bank of fans on the front wing(s) are tiltable.

[0065] Fig. 1H illustrates an alternate arrangement of electric propulsion units, consistent with embodiments of the present disclosure. Referring to Fig. 1H, the aircraft shown in the figure may be a top plan view of an exemplary aircraft. The aircraft may include six electric propulsion systems distributed across the aircraft. In some embodiments, a distribution of electric propulsion systems may include a first set of four electric propulsion systems 185, 186, 187, and 188 coplanar in a first plane and a second set of two electric propulsion systems 189 and 190 coplanar in a second plane. In some embodiments, the first set of electric propulsion systems 185, 186, 187, and 188 may be operatively connected to tilt propellers and second set of electric propulsion systems 189 and 190 may be operatively connected to lift propellers. In other embodiments, the first set of electric propulsion systems 185, 186, 187, and 188 and the second set of aft electric propulsion systems 189 and 190 may all be operatively connected to tilt propellers.

[0066] Fig. II illustrates an alternate arrangement of electric propulsion units, consistent with embodiments of the present disclosure. Referring to Fig. II, the aircraft shown in the figure may be a top plan view of an exemplary aircraft. The aircraft may include four electric propulsion systems distributed across the aircraft. In some embodiments, a distribution of electric propulsion systems may include four coplanar electric propulsion systems 191, 192, 193, and 194. In some embodiments, all of the electric propulsion systems may be operatively connected to tilt propellers.

[0067] Fig. 1 J illustrates an alternate arrangement of electric propulsion units, consistent with embodiments of the present disclosure. Referring to Fig. 1 J, the aircraft shown in the figure may be a top plan view of an exemplary aircraft (e.g., a VTOL aircraft). The aircraft may include six electric propulsion systems distributed across the aircraft. For example, insome embodiments, the aircraft may include four forward electric propulsion systems 195, 196, 197, and 198 operatively connected to tilt propellers and the two aft electric propulsion systems 199 and 200 operatively connected to lift propellers. In some embodiments, the aircraft may include ten electric propulsion systems distributed across the aircraft. For example, in some embodiments, the aircraft may include six forward electric propulsion systems operatively connected to tilt propellers and the four aft electric propulsion systems operatively connected to lift propellers. In some embodiments, some or all of the aft electric propulsion systems may operatively connected to tilt propellers.

[0068] As shown in Fig. 1 J, in some embodiments, the aircraft may have a flying wing configuration, such as a tailless fixed-wing aircraft with no definite fuselage. In some embodiments, the aircraft may have a flying wing configuration with the fuselage integrated into the wing. In some embodiments, the tilt propellers may rotate in a plane above the body of the aircraft when the tilt propellers operate in a lift configuration.

[0069] Figure 2 illustrates a circuit diagram for a High Voltage Junction Box (HVJB), consistent with embodiments of the present disclosure. HVJB 222 may be electrically connected to the HV loads 210 to provide high voltage power. Specifically, 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 BMU DC / DC allows the BMU 271 (shown in Fig. 3) to be powered off the high voltage battery cells BT1. Power storage element BT1 is connected to each of the HV loads through pre-charge resistor(s) (e.g., resistor Rl, R7) 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, fuses F1-F7 may be one or more of the fuses detailed above with respect to Fig. IB. For example, in some embodiments, fuses F2 EE1, F3 EE2, and F4.1 or F4.2 Cross link may correspond to fuses 109, 111, and 131 detailed in Fig. ID.

[0070] Fuse Fl may be a pack fuse to disconnect the failing battery pack 120 from the rest of the HVPS system (e.g., based on detecting a crash and severed cut loop, as described below). In some embodiments, Fl may be a pyro-technical fuse.

[0071] Upon failure of a battery pack 120, pyro-technical fuse Fl may receive a signal (e.g. from the associated battery management system) and blow fuse Fl. Therefore, further damage to other equipment (e.g. electric engines, EPUs, connected battery packs) can be avoided. Further, electric engines 110 will still receive power from the paired battery packs 120 within the battery pack unit. For example, upon a battery pack failure, battery pack 1pyro-technical fuse Fl may be blown, but electric engines 1, 12, 4, and 9 may still receive power from battery pack 4. Further, in some embodiments (as described below), upon detecting an emergency response fuses Fl, F2, F3, F4.1, F4.2, and / or F6 may be blown (e.g., by a battery management system). In some embodiments (as described below), upon detecting an emergency response the contactors KI and / or K4 may be opened (e.g., by a battery management system).

[0072] Figure 3A illustrates a diagram of a High Voltage Junction Box 222 (HVJB), including a battery management system, consistent with embodiments of the present disclosure. In some embodiments, each battery pack 120 contains HV distribution 211, a Battery Management System (BMS 270), and a Pyro-fuse Redundant Trigger board (PRT 280) housed within the HVJB 122. BMS 270 may include one or more processors, microprocessors, controllers, and / or memory storage devices and may be configured to monitor voltages, temperatures, currents, and isolation resistances. BMS 270 may control battery pack contactors and fuses (e.g., shown in HV distribution 211) to protect against fault conditions. As further detailed below, BMS 270 may communicate with various systems within and outside HVJB 222. BMS 270 may include a Battery Management Unit (BMU 271) (e.g., one or more controllers, processors, microprocessors, and / or memory storage devices) which may receive voltage, current, resistance, and temperature sensing signals from cell stack assembly 224 and / or the HV distribution 211. BMS 270 may further include Cell Management Units (CMUs) 272 to monitor the voltages of each set of 7 parallel cells (i.e., a 1S-7P cell group) connected in series in a 14S-7P cell block. CMUs 272 may also be used to monitor a 14S-7P cell block’s temperature. CMUs 272 obtain measurements for all the cell groups in battery pack 120 and communicate the measurements to BMU 271.

[0073] BMU 271 may monitor output current for each of the connected loads. BMU 271 may be internally powered by the battery cell stack assembly 224 and continuously monitor the state of the battery. By monitoring battery pack 120, cell block, and cell group parameters, BMU 271 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 BMU 271 allows for fault detection and isolation through built-in-tests (BIT). In addition, BMU 271 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 battery pack 120. BMU 271 also controls and monitors bus pre-charging, provides fuse and contactor commands, and communicates with various systems within and outside HVJB 222.

[0074] BMU 271 also detects aircraft state information, such as acceleration information. For example, in some embodiments, HVJB 222 may include an accelerometer 291 (e.g., mounted within the battery pack enclosure) and BMU 271 may receive accelerometer measurements and determine whether there has been a crash. In some embodiments, BMU 271 may receive acceleration information from one or more accelerometers inside and / or outside HVJB 222, an inertial measurement system, and / or FCS 230. For example, FCS 230 may provide measurements from an accelerometer and / or an indication that the aircraft has crashed (e.g., based acceleration measurements). In some embodiments, as further described below, BMU 271 may receive information from a triaxial accelerometer and determine a crash based on comparing the acceleration values along one or more axes to one or more corresponding thresholds and / or time periods. In some embodiments, BMU 271 may receive information from a single axis accelerometer. In some embodiments, BMU 271 may receive acceleration information from a single axis accelerometer oriented along an axis that corresponds to a common crash impact angle (e.g., determined through modeling or experimental data). In some embodiments, BMU 271 may determine a crash using the same accelerometer and / or threshold(s) as an emergency location transmitter device which emits a distress signal in the event of a plane crash to facilitate location of the plane.

[0075] In some embodiments, BMU 271 may receive other aircraft state information such as sensor measurements that measure deformation of an aircraft body and / or tracks aircraft movement. In some embodiments, BMU 271 may determine whether these sensor measurements indicate a crash (e.g., by comparing to one or more threshold values).

[0076] In some embodiments, as further detailed below, BMU 271 may further receive information on a flight mode of the aircraft from sensors and / or FCS 230. In some embodiments, a flight mode may correspond to a pilot selection of an intended mode of operation (e.g., in flight or on ground). In some embodiments, a flight mode may be based on one or more sensor detecting whether the aircraft is operating in flight or on the ground.

[0077] BMU 271 includes terminals for low voltage cut loop 290 and monitors a current and / or voltage across cut loop 290. As further detailed below, low voltage cut loop 290 is a low voltage circuit that provides a means for first responders to disconnect battery packs 120 when responding to an emergency situation. As shown above with respect to Fig. 2, in some embodiments, BMU 271 has a dedicated DC / DC converter in battery pack 120 allowing for powering of BMU 271 and cut loop 290, even if there is a loss of power to low voltage power system (LVPS) 240.

[0078] HV distribution 211 in HVJB 222 may contain HV contactors 212 and a combination of active and passive fuses (e.g., pyrotechnical fuses 213 and fuses 214) to protect against overcurrent and short-circuit conditions. In some embodiments, contactors 212 may correspond to one or more of switching devices K1-K7 (e.g. HV contactors) detailed in Fig.2. Similarly, pyrotechnical fuses 213 and fuses 214, may correspond to one or more fuses Fl- F8 detailed in Fig. 2. In some embodiments, a pyro-fuse redundant trigger board (PRT 280) may be located within HVJB 222. BMS 270 may detect a failure event (or emergency response) and send command signals to PRT 280 for a corresponding pyro fuse driver to blow a fuse. In some embodiments, HVJB 122 may further provide a redundant active trigger board configured to enable the pyro fuse driver to activate one or more pyrotechnical fuses when BMS 270 fails to enable the pyro fuse driver.

[0079] In some embodiments, battery packs 120 may be in communication with each other (e.g. through BMS 270 and associated BMU 271). In some embodiments, upon detection of an emergency response (as further detailed below), one battery pack 120 may communicate with other battery packs to disconnect high voltage power (e.g., blow one or more fuses and / or open one or more contactors). The communication between battery packs may be through direct communication lines and / or through a central control unit (e.g., including one or more processors and / or microprocessors)

[0080] Control MCU (CCU 263) in charge port assembly 262 may interface with the external battery charger and communicate with 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, CCU 263 may be a single PCB A with one microcontroller that manages overall power delivery to each battery pack 120 when charging. In some embodiments, Control MCU 263 may handle communication between battery packs 120.

[0081] In some embodiments, BMU 271 may be exclusively powered by cell stack 224, while in other embodiments BMU 271 and / or other associated components may be powered by a separate low voltage power source (LVPS) 240 (e.g., a low voltage battery). In some embodiments, battery pack 120 is directly connected to aircraft switches 250 (e.g., associated with the FCS 230), while in other embodiments battery pack 120 is not directly connected to aircraft switches 250.

[0082] Figure 3B illustrates a central battery management system, consistent with embodiments of the present disclosure. In some embodiments, the aircraft may include a central battery management system 270 (e.g., including a battery management unit 271). In some embodiments, central battery management system 270 may perform some or all of thesame functions as battery management system 270 shown in Fig. 3 A for multiple battery packs 120. For example, central battery management system 270 may receive voltage, current, resistance, and / or temperature information from each battery pack 120 and / or associated high voltage wiring connected to battery packs 120. As described above, the central battery management system 270 may control the contactors 212, pyro fuses 213, fuses 214, based on the received information.

[0083] Further, central battery management system (BMS) 270 may receive aircraft state information from one or more accelerometers 291 and / or FCS 230. For example, central battery management system 270 may receive acceleration information from one or more accelerometers 291 and / or FCS 230. While three accelerometers are shown, central BMS 270 may receive acceleration information from any number of accelerometers 291 (e.g., 1, 2, 3, or 4 accelerometers 291). In some embodiments, accelerometers 291 may be spaced throughout the aircraft in positions that may best allow them to detect a crash (e.g., as determined by crash simulations). In some embodiments (as further described below), BMS 270 may receive a flight mode from FCS 230 and / or sensors that detect a flight mode. Further, BMS 270 may include a separate LV power source 310 to provide power to one or more control unit(s) of the BMS 270. In some embodiments, BMS 270 may include one control unit (e.g., BMU 271), while in other embodiments BMS 270 may include multiple controllers (e.g., control units 301 and 302, each including one or more processors, microprocessors, and / or memory storage devices) to provide redundancy for detecting condition(s) and controlling contactors 212, pyro fuses 213, and / or fuses 214. For example, as shown by the two switching devices 312 between the battery pack(s) 120 and power source 310, both control units (e.g., 301 and 302) must detect a condition and close a switch to active a pack disconnection device (e.g., blow a fuse). Switching devices 312 may include any switching device capable of enabling and / or disabling power to a disconnection device (e.g., a relay, transistor, contactor etc.). While these switching devices 312 are shown as part of BMS 270, in other embodiments, switching devices 312 may be included in the respective battery packs 120. In some embodiments, agreement between control units (301 and 302) to activate a pack disconnection device may be established through a combination of logic gates. In some embodiments, agreement between control units (301 and 302) to activate a pack disconnection device may be established by communication between the control units and associated programming to determine an agreement. While two control units are shown, the central BMS 270 may include a different number of control units (e.g., three, four, five etc.).

[0084] As further described below, in some embodiments both Control Units 301 and 302 must detect aircraft state information indicates a crash (e.g., acceleration exceeds a threshold) and a cut loop 290 has been cut to disconnect a battery pack 120. In some embodiments, both Control Units 301 and 302 must detect aircraft state information indicates a crash and one control unit must detect a cut loop 290 has been cut in order to disconnect a battery pack 120. While one cut loop 290 is shown, in some embodiments, there may be multiple cut loops 290 which control the disconnection of one or more battery packs 120. In some embodiments, cut loop(s) may be associated with the disconnection of one battery pack or a subgroup of battery packs, while in other embodiments all cut loop(s) may trigger the disconnection of all battery packs 120.

[0085] Figure 4A illustrates a flow chart for detecting an emergency response, consistent with embodiments of the present disclosure. In some embodiments, this process is performed by one or more processors in a battery management system 270. For example, this process may be performed by a BMS 270 in each battery pack (e.g., as shown in Fig. 3 A) or by a central BMS 270 (as shown in Fig. 3B). In some embodiments, this process is performed by one or more processors in a flight control system. In some embodiments, different steps are performed by different processors and / or systems.

[0086] At Step 401, the one or more processors receive acceleration information. Acceleration information may be received directly from sensors (e.g. an accelerometer in a battery pack 120), an inertial measurement system, and / or a flight control system of the aircraft. In some embodiments, each battery pack 120 has a triaxial accelerometer. For example, each battery pack 120 may include a triaxial accelerometer mounted within the battery pack’s enclosure. In some embodiments, the acceleration information is in the form of measured data, while in other embodiments the acceleration information is an estimated acceleration based on measured data. In some embodiments, the acceleration information received by the battery management system 270 is filtered to attenuate the impacts of aircraft vibrations (e.g., propeller vibrations). Acceleration information indicates a total acceleration and / or an acceleration along a first, second, and / or third axis. In some embodiments, one or more accelerometers may be mounted throughout the aircraft. For example, the one more more accelerometers may be mounted in the tail of the aircraft.

[0087] At Step 402, the one or more processors receive continuity information (e.g. from BMU 271) indicating whether or not a low voltage emergency cut loop (e.g., cut loop 290) has been cut. In some embodiments, the one or more processors determine whether or not a low voltage emergency cut loop 290 has been cut. For example, the one or more processorsdetermine that a cut loop 290 has been cut based on detecting a loss of current and / or voltage across the cut loop 290. Continuity information indicating the cut loop 290 has been cut may be the result of a first responder cutting the loop or a severe crash that results in the cut loop 290 being cut.

[0088] The information gathered in Steps 401 and 402 is received sequentially or simultaneously. Further, in some embodiments, the received information includes a time stamp indicating when it was collected. In other embodiments, the one or more processors assign a time based on when the information was received.

[0089] At Step 403, the one or more processors receive a flight mode of the aircraft. In some embodiments, the one or more processors receive the flight mode from a flight control system (FCS) of the aircraft. Further, in some embodiments, the flight mode indicates a pilot selected mode of aircraft operation, such as an “off mode” where the pilot has powered off the aircraft, “service mode” where the pilot has enabled the aircraft to receive service, “ground mode” where the pilot has indicated the aircraft is on the ground, and / or “fly mode” where the pilot has indicated the aircraft is in flight and / or will be in flight shortly. The “off mode” and “service mode” may have the same effect as a “ground mode” when determining an emergency response. In some embodiments, the flight mode may simply indicate a “ground mode” where the pilot has indicated the aircraft is on the ground and a “fly mode” where the pilot has indicated the aircraft is in flight and / or will be in flight shortly.

[0090] In some embodiments, the flight mode may be determined (e.g., by FCS) using one or more sensors, such as a pitot tube to measure airspeed, landing gear switch(es), sensors detecting flight element states (e.g., state of control surface(s)), altitude sensors, and / or wheel sensors (e.g., weight on wheels switch) to detect deployed aircraft wheels. In some embodiments, a flight mode may be detected using a combination of sensor inputs. For example, “fly mode” may be detected based on an airspeed exceeding a threshold level (e.g., currently and / or within a past threshold period of time) and landing gear not being engaged. For example, “ground mode” may be detected based on airspeed being below a threshold level and landing gear being engaged. An airspeed threshold may be set based on experimental data and / or modeling aircraft flight conditions to determine at what airspeeds the aircraft is in flight.

[0091] At Step 404, the one or more processors determine whether there has been a crash based on the received acceleration information and the received continuity information. In some embodiments, the one or more processors determine a crash based on a first condition being met and then a second condition being met. The first condition includes determiningacceleration information indicates a crash. In some embodiments, the one or more processors determine acceleration information indicates a crash when acceleration exceeds a threshold and / or exceeds a threshold for a set period of time. An acceleration threshold may be determined based on experimental data and / or modeling aircraft acceleration during a crash. In some embodiments, the set period of time may be determined to avoid erroneous crash detection (e.g., based on a spike in acceleration not associated with a crash). “Acceleration” refers to an absolute value of negative acceleration corresponding to a decrease in velocity.

[0092] In some embodiments, a total acceleration is compared to a threshold to determine whether the acceleration indicates a crash. For example, acceleration information may be received from a single axis accelerometer oriented along an axis that corresponds to a common crash impact angle (e.g., determined through modeling or experimental data) and may be compared to a threshold. In some embodiments, acceleration across a first, second, and / or third axis (e.g., longitudinal, vertical, and / or lateral) is compared to a first, second, and / or third threshold to determine whether the acceleration indicates a crash. For example, in some embodiments, a crash is determined when a longitudinal acceleration exceeds a first threshold, such as in a “winge-borne” crash scenario when the aircraft’s lift is primarily supported by the airflow over the wings of the aircraft. Further, a crash is determined when a vertical acceleration exceeds a second threshold, such as in a “thrust-borne” crash scenario when an aircraft’s lift is primarily supported by the electric propulsion units (e.g., electric engines and propellers). Further, a crash is determined when a lateral acceleration exceeds a third threshold. In some embodiments, the thresholds for detecting crashes may be determined based on modeling the impacts to acceleration during different crash scenarios. In some embodiments, exceeding one or more combination(s) of different thresholds may indicate a crash. For example, a combination of a longitudinal acceleration above a fourth threshold and a vertical acceleration above a fifth threshold may indicate a crash. As described above, the different combinations of thresholds may be set according to modeling or experimental data.

[0093] In some embodiments, other aircraft state information, instead of or in addition to aircraft acceleration, may be used for the first condition to indicate a crash. For example, the one or more processors may determine there is a crash based on sensor measurements indicating aircraft deformation and / or tracking aircraft movement.

[0094] The second condition includes determining the received continuity information indicates the cut loop 290 is cut (e.g., no current and / or voltage in the cut loop 290). In some embodiments, the one or more processors determine a crash when the accelerationinformation indicates a crash at an earlier time than the continuity information indicates a cut loop.

[0095] In some embodiments, in addition to determining a crash based on acceleration information indicating a crash and then continuity information indicating a cut loop, the one or more processors may determine a crash based on acceleration information indicating a crash within a threshold of time after the continuity information indicates a cut loop. Therefore, the one or more processors may detect a crash and disconnect power in a situation where the initial impact of the crash severs the cut loop before the acceleration information indicates a crash. For example, the one or more processors may detect a crash when the tail of the aircraft impacts the ground first, immediately severing the cut loop. The threshold may be set based on experimental data and / or modeling aircraft crash conditions to determine a period of delay between the continuity information indicating a cut loop and acceleration information indicating a crash. In some embodiments, the threshold period of time may be set to 5 seconds or less (e.g., 5, 4, 3, 2, 1 seconds, or a value less than a second).

[0096] By only determining a crash when the acceleration information indicates a crash before the continuity information indicates a cut loop (or within a threshold of period of time after), the one or more processors ensure that the battery packs are not disconnected based on an open-circuit failure not associated with a crash.

[0097] At Step 405, based on the one or more processors determining a crash, the one or more processors send a command to disconnect at least one battery pack 120 from high voltage power. For example, the one or more processors send a command to blow a fuse (e.g., battery pack fuse Fl, electric engine fuses F2, F3, and / or cross link fuse(s) F4, F6), open a contactor, open a relay, and / or open another switching device to disconnect one or more battery packs 120. In some embodiments, the one or more processors may only command a disconnect of battery pack 120 connected to the cut loop 290 that is cut, while in other embodiments the one or more processors command a disconnect of additional battery packs 120 or all battery packs 120. In some embodiments, the one or more processors command the disconnect of all battery packs 120 that have a high voltage wiring connection (e.g., a cross-link) with battery pack 120 associated with the cut loop 290. For example, referencing Fig. 1 A, based on determining that an emergency responder cut loop 1 associated with battery pack 1, the one or more processors may blow the pyrotechnical fuses associated with battery packs 1 and 4.

[0098] At Step 406, the one or more processors determine whether the acceleration information indicates a false crash based on the received acceleration information, thecontinuity status, and the flight mode. In some embodiments, the one or more processors determines a crash based on a first condition being met, a second condition being met, and a third condition being met. The first condition includes determining acceleration information indicates a crash (same first condition of Step 404). The second condition includes determining the received continuity information does not indicate a cut loop 290 is cut. The third condition includes determining the flight mode indicates a ground mode. For example, in some embodiments, the received flight mode includes a pilot selection of “ground mode” on a pilot interface. In some embodiments, sensor information may indicate the aircraft is in a ground modem. For example, a sensor for wheel deployment may indicate an aircraft is in ground mode.

[0099] The one or more processors determine the acceleration information indicates a false crash when the acceleration information indicate a crash, the continuity information indicates no cut loop, and the flight mode indicates a ground mode. This may occur upon a hard landing where acceleration was great enough to meet the condition(s) for indicating a crash.

[0100] At Step 407, based on determining acceleration information indicates a false crash, the one or more more processors will reset condition 1 for Steps 404, 406, and 408. Further, the one or more processors will de-active an alert initiated in Step 409. Further, new acceleration information will be gathered at Step 401.

[0101] At Step 408, the one or more processors determine whether there is a potential crash without a response based on the received acceleration information, the continuity status, and the flight mode. In some embodiments, the one or more processors determine a potential crash without response based on a first condition being met, a second condition being met, and a third condition being met. The first condition includes determining acceleration information indicates a crash (same first condition of Step 404). The second condition includes determining the received continuity information does not indicate a cut loop (no first responder yet). The third condition includes determining the flight mode does not indicate a ground mode. For example, in some embodiments, the flight mode does not indicate a ground mode when the received flight mode indicates a pilot selection of a “fly mode” on a pilot interface. In some embodiments, flight mode does not indicate a ground mode when the aircraft is in flight and / or has been in flight within a threshold period of time.

[0102] The one or more processors determine a potential crash without response when the acceleration information indicate a crash, the continuity status indicates no cut loop, and the indicated flight mode is not ground mode.

[0103] At Step 409, based on determining a potential crash without response, the processor sends a command to issue an alert (e.g., send a signal, provide a notice, provide a warning etc.). For example, the command provides an alert on a display screen (e.g., a display screen mounted in the tail of the aircraft and / or a flight avionics display), turns on a light (e.g., a light on the battery pack), and / or activates a sound (e.g., a buzzer in the battery pack and / or spoken notification). In some embodiments, the aircraft includes markings and / or signs to indicate the meaning of the alert to a first responder. For example, a sign may indicate that the high voltage power is still active when the buzzing sound is present. In some embodiments, the alert will continue until deactivated by a manual input (e.g., on battery pack and / or display screen) and / or by a false crash being determined at Step 406.

[0104] Figure 4B illustrates a flow chart for detecting an emergency response, consistent with embodiments of the present disclosure. In some embodiments, this process is performed by one or more processors in each battery management system 270 of battery packs 120. In some embodiments, this process is performed by one or more processors in a flight control system. In some embodiments, different steps are performed by different processors and / or systems.

[0105] Steps 401-407 are described above with reference to Fig. 4A. At Step 410, the one or more processors determine whether there is a potential crash without the acceleration information indicating a crash (e.g., first condition Step 404 not met), based on the received acceleration information, the continuity information, and the flight mode. In some embodiments, the one or more processors determine a potential crash without acceleration detection based on a first condition being met, a second condition being met, and a third condition being met. The first condition includes determining acceleration information does not indicate a crash (e.g., does not meet first condition of Step 404 requirements). The second condition includes determining the received continuity information indicates a cut loop 290 is cut (e.g., via a first responder). The third condition includes determining the flight mode does not indicate ground mode. For example, in some embodiments, the received flight mode is based on a pilot selection of a “fly mode” on a pilot interface. In some embodiments, the received flight mode is based on one or more sensor measurements indicating the aircraft is in flight and / or has been in flight within a threshold period of time.

[0106] At Step 411, based on determining the potential crash without acceleration detection, the one or more processors sends a command to issue an alert. For example, the command provides an alert on a display screen (e.g., a display screen mounted in the tail of the aircraft and / or a flight avionics display), turns on a light (e.g., a light on the battery pack), and / oractivates a sound (e.g., a buzzer in the battery pack and / or spoken notification). In some embodiments, the aircraft includes markings and / or signs to indicate the meaning of the alert to a first responder. For example, a sign may indicate that the high voltage power is still active when the buzzing sound is present.

[0107] In some embodiments, the one or more processors may additionally or alternatively transmit a maintenance message based on determining the potential crash without acceleration detection. In some embodiments, a maintenance message may be sent to flight deck avionics. For example, one or more processors may transmit a maintenance message indicating the severed cut loop to the FCS 230. In some embodiments, the maintenance message may indicate information on the severed cut loop, such as an associated battery pack and / or routing location. The FCS 230 may provide the maintenance message through text and / or images on a display screen, through turning on light(s), and / or through activating sounds (e.g., a spoken notification or noise alert). A “maintenance message” may refer to any of the provided information.

[0108] In some embodiments, the one or more processors may include both Steps 408-409 and Steps 410-411. For example, a first alert may indicate a potential crash without a response and a second alert may indicate a potential crash without acceleration detection. In some embodiments, the alerts provided by the one or more processors are the same, while in other embodiments they are different.

[0109] Figures 5A-5C illustrate areas where the high voltage wiring is located on the aircraft, consistent with embodiments of the present disclosure. In Fig. 5A, the dashed marking shows high voltage wiring area 502. As shown, the high voltage wiring is run outside of the cabin area 501 where occupants of the aircraft are located. There is no high voltage wiring in the aircraft doors, exterior cabin walls, or ceiling area above the cabin. Therefore, a first responder may safely cut into the cabin area to extract occupants. Further, in some embodiments, there is no high voltage wiring in the aircraft landing gear and the aircraft tail, reducing risk of high voltage wiring exposure upon a crash that damages these aircraft elements. In some embodiments, the high voltage wiring is separated from cabin area 501 by a bulkhead and / or other physical barrier that protects occupants and rescuers from high voltage wiring. A bulkhead 504 may separate the cabin area from high voltage wiring area 502. For example, a structural bulkhead 504 that carries wing loads and / or landing gear loads into the body of the aircraft, and furnishings on bulkhead 504, may separate cabin area 501 from high voltage area 502. In some embodiments, the wing of the aircraft including high voltage wiring is aft of the cabin area. In other embodiments the wing of the aircraftincluding high voltage wiring is above the cabin area and the high voltage wiring does not penetrate into cabin area 501. Further, the separation between the cabin area 501 and high voltage area 502 may be indicated by markings, text, and / or signs on the skin of the aircraft. As shown, first responders can safely cut through any part of the doors, pillars, floor, and through the part of the cabin ceiling not blocked by the wing.

[0110] In some embodiments, as shown in Fig. 5B, the high voltage wiring may run between battery packs 120 and the electric engines within high voltage wiring area 502. In some embodiments, all equipment with high voltage wiring connections (e.g., heating / cooling equipment and / or other auxiliary loads) may be included in the high voltage wiring area 502. As further detailed below, low voltage emergency cut loop 290 may run from battery packs 120 in the high voltage wiring area 502 through the rear of the aircraft 503. In some embodiments, low voltage emergency cut loop 290 may run from battery packs 120 in high voltage wiring area 502 through cabin area 501. Therefore, a first responder may cut one or more loops to de-energize battery packs 120 without risking cutting into an energize high voltage wiring, thereby increasing safety. Further, the first responder may cut one or more cut loops 290 to de-energize battery packs 120 while avoiding aircraft propellers that may still be spinning post-crash, thereby increasing safety.[H l] As shown in Fig. 5C, in some embodiments, all battery packs 120 supplying high voltage power are located within the aircraft wings. For example, each battery pack 120 may be located in the wing box(es) 512, between the spars 508 and the ribs 512. The high voltage wiring (as represented by dashed lines 516) may run laterally across the wing of the aircraft through holes in the spars 510, connecting battery packs 120 and electric propulsion units 110. In some embodiments, equipment requiring high voltage power (e.g., air conditioning etc.) may be installed aft of the cabin (e.g., aft of bulkhead 504). The high voltage wiring may be run though the aft and / or bottom of the aircraft wing to feed the equipment in a manner that avoids the cabin area 501.

[0112] Figs. 6A-6B illustrate plan view diagrams for routing low voltage emergency cut loop wiring through the tail of an aircraft, consistent with embodiments of the present disclosure. As detailed above, in some embodiments, each low voltage emergency cut loop 290 may be connected to a single battery pack 120. Therefore, if an aircraft has six battery packs 120, six cut loops 290 may be routed from battery packs 120. In some embodiments, one or more battery packs 120 may share a cut loop 290. Therefore, in some embodiments, the number of cut loops 290 is less than the number of battery packs 120. In some embodiments, sharing a cut loop may involve physically connecting cut loop 290 to bothbattery packs 120, so that either battery pack can detect a loss in continuity (e.g., current or voltage). In some embodiments, sharing a cut loop 290 may involve physically connecting the cut loop 190 to a battery pack 120 and that battery pack 120 communicating with one or more other battery packs 120 when the cut loop 290 has been cut. In some embodiments, cross-linked battery packs 120 may share a cut loop 290. In some embodiments, the cut loops 290 are routed from battery packs 120 located in the wings, or elsewhere, to the tail of the plane. In some embodiments, the cut loops 290 include tags and / or painting to indicate their purpose and / or effect.

[0113] As described above, this routing ensures that cut loops 290 are accessible to be cut in rear of the aircraft (e.g., rear of aircraft 503) away from the high voltage wiring running between the batteries, electric engines, and other aircraft devices located towards the front of the aircraft (e.g., high voltage wiring area 502).

[0114] As shown in Fig. 6A, in some embodiments, each cut loop 290 is routed separately. As shown in Fig. 6B, in some embodiments, one or more cut loops 290 are combined (e.g. in a bundle). In some embodiments, cut loops 290 of battery packs 120 with a high voltage wiring connection (e.g. cross-link) are combined. For example, referencing Fig. 1 A, battery packs 1 and 4 are cross-linked and the cut loops associated with these battery packs are bundled together. The number of bundled cut loops may depend on the number of battery packs in a battery pack unit of cross-linked battery packs. For example, as described above, a battery pack unit may include two, three, four, or any number of battery packs, and the number of cut loops in a bundle may equal the number of battery packs in the battery pack unit.

[0115] In some embodiments, cut loop(s) connected to each battery pack unit may be spaced apart from cut loop(s) connected to other battery pack units to avoid a single point of failure affecting multiple battery pack units. For example, with reference to Fig. 1 A, cut loop(s) associated with battery packs 1 and 4 may be spaced separately from cut loop(s) associated with battery packs 3 and 6 and separately from cut loop(s) associated with battery packs 2 and 5. In some embodiments, the spacing maintained between cut loop(s) may be at least a few inches.

[0116] Figure 6C illustrates a profile view diagram for routing cut loop wiring through the tail of an aircraft, consistent with embodiments of the present disclosure. As detailed above, cut loops 290 are routed to the tail of the aircraft to increase safety of the emergency responders. Additionally, cut loops 290 may be routed in a manner that allows them to be easily accessed by a first responder. For example, in some embodiments, the cut loops maybe routed towards the perimeter of the aircraft so they are easier to find and cut. In some embodiments, cut loops may be color coded and contain descriptive tags at set intervals to ensure first responders are easily able to identify them.

[0117] The foregoing description has been presented for purposes of illustration. It is not exhaustive and does not limit the invention to the precise forms or embodiments disclosed. Modifications and adaptations of the invention will be apparent to those skilled in the art from consideration of the specification and practice of the disclosed embodiments of the inventions disclosed herein.

[0118] The features and advantages of the disclosure are apparent from the detailed specification, and thus, it is intended that the appended claims cover all systems and methods falling within the true spirit and scope of the disclosure. As used herein, the indefinite articles “a” and “an” mean “one or more.” Similarly, the use of a plural term does not necessarily denote a plurality unless it is unambiguous in the given context. Words such as “and” or “or” mean “and / or” unless specifically directed otherwise. Further, since numerous modifications and variations will readily occur from studying the present disclosure, it is not desired to limit the disclosure to the exact construction and operation illustrated and described, and accordingly, all suitable modifications and equivalents may be resorted to, falling within the scope of the disclosure.

[0119] Other embodiments will be apparent to those skilled in the art from consideration of the specification and practice of the implementations disclosed herein. It is intended that the architectures and circuit arrangements shown in figures are only for illustrative purposes and are not intended to be limited to the specific arrangements and circuit arrangements as described and shown in the figures. It is also intended that the specification and examples be considered as exemplary only, with the true scope and spirit of the invention being indicated by the following claims. The foregoing description has been presented for purposes of illustration. It is not exhaustive and does not limit the invention to the precise forms or embodiments disclosed. Modifications and adaptations of the invention will be apparent to those skilled in the art from consideration of the specification and practice of the disclosed embodiments of the inventions disclosed herein.

[0120] The embodiments may further be described using the following clauses:1. A method of controlling aircraft power distribution, comprising: receiving aircraft state information of an aircraft; detecting whether the aircraft state information indicates a crash; anddisconnecting supply of high voltage power to the aircraft by at least one battery upon detecting both that the aircraft state information indicates a crash and a loss of continuity in at least one low voltage wire. ethod of clause 1, wherein the aircraft state information is acceleration information received from at least one accelerometer. ethod of clause 2, wherein the at least one accelerometer is mounted to the at least one battery or in a tail of the aircraft. ethod of any of clauses 1-3, wherein detecting that the aircraft state information indicates a crash comprises detecting aircraft acceleration exceeds a threshold. ethod of clause 4, wherein detecting that the aircraft state information indicates a crash comprises detecting that aircraft acceleration exceeds the threshold for a set period of time. ethod of any of clauses 1-5, wherein detecting that the aircraft state information indicates a crash comprises: detecting that a longitudinal acceleration exceeds a longitudinal acceleration threshold, or detecting that a vertical acceleration exceeds a vertical acceleration threshold. ethod of any of clauses 1-6, wherein detecting that the aircraft state information indicates a crash comprises detecting that aircraft acceleration exceeds a threshold set to trigger an emergency location transmitter of the aircraft. ethod of any of clauses 1-7, wherein the at least one low voltage wire is connected to the at least one battery. ethod of any of clauses 1-8, wherein the at least one low voltage wire is routed through the tail of the aircraft. method of any of clauses 1-9, wherein disconnecting the supply of the high voltage power comprises at least one of: blowing a fuse of the at least one battery or opening a contactor of the at least one battery. method of any of clauses 1-10, wherein disconnecting the supply of the high voltage power comprises blowing a pyro-technical fuse of the at least one battery. method of any of clauses 1-11, further comprising issuing an alert upon detecting both that the aircraft state information does not indicate a crash and a loss of continuity in the at least one low voltage wire. method of clause 12,wherein issuing an alert upon detecting a loss of continuity in the at least one low voltage wire further comprises: receiving a mode of the aircraft from a flight control method of the aircraft; and issuing the alert upon detecting that the aircraft was set to a fly mode, wherein the fly mode is based on a pilot selection or sensor measurements the indicating aircraft was in flight. method of clause 12 or 13, wherein issuing the alert comprises turning on a light or activating a sound. method of any of clauses 12-14, wherein issuing the alert comprises activating a buzzer. method of any of clauses 1-15, wherein disconnecting the supply of the high voltage power to the aircraft comprises: disconnecting the supply of high voltage power upon detecting that the aircraft state information indicates a crash prior to the loss of continuity in the at least one low voltage wire. method of any of clauses 1-16, wherein disconnecting the supply of the high voltage power to the aircraft further comprises: disconnecting the supply of high voltage power upon detecting that the aircraft state information indicates a crash within a threshold of time after the loss of continuity in the at least one low voltage wire. method of any of clauses 1-17, further comprising: determining a false crash detection upon: detecting aircraft state information indicates a crash, detecting no loss of continuity in the at least one low voltage wire, and detecting that the aircraft is in a ground mode, wherein the ground mode is based on a pilot selection or sensor measurements indicating the aircraft is on the ground; and receiving new aircraft state information upon determining the false crash detection. method of any of clauses 1-18, further comprising: issuing an alert upon detecting the aircraft state information indicates a crash and detecting no loss of continuity in the at least one low voltage wire.method of any of clauses 1-19, further comprising: issuing an alert upon detecting no loss of continuity in the at least one low voltage wire and detecting the aircraft state information indicates a crash; determining a false crash detection upon detecting that the aircraft is in a ground mode, wherein the ground mode is based on a pilot selection or sensor measurements indicating the aircraft is on the ground; and deactivating the alert upon determining the false crash. electrical system comprising at least one processor configured to perform the method of any one of clauses 1-20. aircraft comprising the electrical system of clause 21. electrical system for an aircraft, comprising: a first battery; a second battery; a first low voltage cut loop connected to the first battery and a second low voltage cut loop connected to the second battery, wherein the first low voltage cut loop is spaced apart from the second low voltage cut loop. electrical system of clause 23, wherein the first and second low voltage cut loops are routed through a section of a fuselage of the aircraft free of high voltage wiring. electrical system of clause 23 or 24, wherein the first and second low voltage cut loops are routed through a tail of the aircraft or a cabin of the aircraft. electrical system for an aircraft, comprising: a first battery; a second battery; a communication wire routed between the first battery and the second battery; a first low voltage cut loop connected to the first battery, wherein no low voltage cut loop is connected to the second battery. electrical system of clause 26, wherein the first low voltage cut loop is routed through a section of a fuselage of the aircraft free of high voltage wiring. electrical system of clause 26 or 27, wherein the first low voltage cut loop is routed through a tail of the aircraft or a cabin of the aircraft. electrical system for an aircraft, comprising:a first paired battery pack unit comprising a first battery and a second battery, the first battery electrically connected to the second battery via a first high voltage bus; a second paired battery pack unit comprising a third battery and a fourth battery, the third battery electrically connected to the fourth battery via a second high voltage bus; wherein the first high voltage bus and second high voltage bus are electrically separate from one another; and a first low voltage cut loop connected to the first battery and a second low voltage cut loop connected to the third battery. electrical system of clause 29, wherein the first low voltage cut loop is routed through a section of a fuselage of the aircraft free of high voltage wiring. electrical system of clause 29 or 30, wherein the first low voltage cut loop is routed through a tail of the aircraft or a cabin of the aircraft. electrical system of any of clauses 29-31, wherein the first low voltage cut loop is spaced apart from the second low voltage cut loop. electrical system of any of clauses 29-32, further comprising: a third low voltage cut loop connected to the second battery and a fourth low voltage cut loop connected to the fourth battery. electrical system of clause 33, wherein the first and second low voltage cut loops are bundled in a first bundle and the third and fourth low voltage cut loops are bundled in a second bundle; and wherein the first bundle is spaced apart from the second bundle. aircraft, comprising: a fuselage; a wing connected to the fuselage; a cabin area section of the fuselage; high voltage wiring; battery packs installed in the wing; wherein walls and floor of the cabin area section are free of the high voltage wiring; and wherein the wing includes the high voltage wiring. aircraft of clause 35, wherein the high voltage wiring runs between the battery packs.aircraft of clause 35 or 36, wherein the high voltage wiring runs laterally between the fore and aft spars of the wing. aircraft of any of clauses 35-37, wherein the high voltage wiring runs through ribs of the wing. aircraft of any of clauses 35-38, wherein a ceiling of the cabin area section is free of high voltage wiring. aircraft of any of clauses 35-38, wherein a ceiling of the cabin area section that does not include the wing is free of high voltage wiring. aircraft of any of clauses 35-40, wherein a low voltage cut loop extends from at least one of the battery packs through the cabin area section free of high voltage wiring. aircraft of any of clauses 35-41, wherein a tail section of the aircraft is free of high voltage wiring. aircraft of clause 42, wherein a low voltage cut loop extends from at least one of the battery packs through the tail section of the aircraft free of high voltage wiring.

Claims

CLAIMS:

1. A method of controlling aircraft power distribution, comprising: receiving aircraft state information of an aircraft; detecting whether the aircraft state information indicates a crash; and disconnecting supply of high voltage power to the aircraft by at least one battery upon detecting both that the aircraft state information indicates a crash and a loss of continuity in at least one low voltage wire.

2. The method of claim 1, wherein the aircraft state information is acceleration information received from at least one accelerometer.

3. The method of claim 2, wherein the at least one accelerometer is mounted to the at least one battery or in a tail of the aircraft.

4. The method of any of claims 1-3, wherein detecting that the aircraft state information indicates a crash comprises detecting aircraft acceleration exceeds a threshold.

5. The method of claim 4, wherein detecting that the aircraft state information indicates a crash comprises detecting that aircraft acceleration exceeds the threshold for a set period of time.

6. The method of any of claims 1-5, wherein detecting that the aircraft state information indicates a crash comprises: detecting that a longitudinal acceleration exceeds a longitudinal acceleration threshold, or detecting that a vertical acceleration exceeds a vertical acceleration threshold.

7. The method of any of claims 1-6, wherein detecting that the aircraft state information indicates a crash comprises detecting that aircraft acceleration exceeds a threshold set to trigger an emergency location transmitter of the aircraft.

8. The method of any of claims 1-7, wherein the at least one low voltage wire is connected to the at least one battery.

9. The method of any of claims 1-8, wherein the at least one low voltage wire is routed through the tail of the aircraft.

10. The method of any of claims 1-9, wherein disconnecting the supply of the high voltage power comprises at least one of: blowing a fuse of the at least one battery or opening a contactor of the at least one battery.

11. The method of any of claims 1-10, wherein disconnecting the supply of the high voltage power comprises blowing a pyro-technical fuse of the at least one battery.

12. The method of any of claims 1-11, further comprisingissuing an alert upon detecting both that the aircraft state information does not indicate a crash and a loss of continuity in the at least one low voltage wire.

13. The method of claim 12, wherein issuing an alert upon detecting a loss of continuity in the at least one low voltage wire further comprises: receiving a mode of the aircraft from a flight control method of the aircraft; and issuing the alert upon detecting that the aircraft was set to a fly mode, wherein the fly mode is based on a pilot selection or sensor measurements the indicating aircraft was in flight.

14. The method of claim 12 or 13, wherein issuing the alert comprises turning on a light or activating a sound.

15. The method of any of claims 12-14, wherein issuing the alert comprises activating a buzzer.

16. The method of any of claims 1-15, wherein disconnecting the supply of the high voltage power to the aircraft comprises: disconnecting the supply of high voltage power upon detecting that the aircraft state information indicates a crash prior to the loss of continuity in the at least one low voltage wire.

17. The method of any of claims 1-16, wherein disconnecting the supply of the high voltage power to the aircraft further comprises: disconnecting the supply of high voltage power upon detecting that the aircraft state information indicates a crash within a threshold of time after the loss of continuity in the at least one low voltage wire.

18. The method of any of claims 1-17, further comprising: determining a false crash detection upon: detecting aircraft state information indicates a crash, detecting no loss of continuity in the at least one low voltage wire, and detecting that the aircraft is in a ground mode, wherein the ground mode is based on a pilot selection or sensor measurements indicating the aircraft is on the ground; andreceiving new aircraft state information upon determining the false crash detection.

19. The method of any of claims 1-18, further comprising: issuing an alert upon detecting the aircraft state information indicates a crash and detecting no loss of continuity in the at least one low voltage wire.

20. The method of any of claims 1-19, further comprising: issuing an alert upon detecting no loss of continuity in the at least one low voltage wire and detecting the aircraft state information indicates a crash; determining a false crash detection upon detecting that the aircraft is in a ground mode, wherein the ground mode is based on a pilot selection or sensor measurements indicating the aircraft is on the ground; and deactivating the alert upon determining the false crash.

21. An electrical system comprising at least one processor configured to perform the method of any one of claims 1-20.

22. An aircraft comprising the electrical system of claim 21.

23. An electrical system for an aircraft, comprising: a first battery; a second battery; a first low voltage cut loop connected to the first battery and a second low voltage cut loop connected to the second battery, wherein the first low voltage cut loop is spaced apart from the second low voltage cut loop.

24. The electrical system of claim 23, wherein the first and second low voltage cut loops are routed through a section of a fuselage of the aircraft free of high voltage wiring.

25. The electrical system of claim 23 or 24, wherein the first and second low voltage cut loops are routed through a tail of the aircraft or a cabin of the aircraft.

26. An electrical system for an aircraft, comprising: a first battery; a second battery; a communication wire routed between the first battery and the second battery; a first low voltage cut loop connected to the first battery, wherein no low voltage cut loop is connected to the second battery.

27. The electrical system of claim 26, wherein the first low voltage cut loop is routed through a section of a fuselage of the aircraft free of high voltage wiring.

28. The electrical system of claim 26 or 27, wherein the first low voltage cut loop is routed through a tail of the aircraft or a cabin of the aircraft.

29. An electrical system for an aircraft, comprising: a first paired battery pack unit comprising a first battery and a second battery, the first battery electrically connected to the second battery via a first high voltage bus; a second paired battery pack unit comprising a third battery and a fourth battery, the third battery electrically connected to the fourth battery via a second high voltage bus; wherein the first high voltage bus and second high voltage bus are electrically separate from one another; and a first low voltage cut loop connected to the first battery and a second low voltage cut loop connected to the third battery.

30. The electrical system of claim 29, wherein the first low voltage cut loop is routed through a section of a fuselage of the aircraft free of high voltage wiring.

31. The electrical system of claim 29 or 30, wherein the first low voltage cut loop is routed through a tail of the aircraft or a cabin of the aircraft.

32. The electrical system of any of claims 29-31, wherein the first low voltage cut loop is spaced apart from the second low voltage cut loop.

33. The electrical system of any of claims 29-32, further comprising: a third low voltage cut loop connected to the second battery and a fourth low voltage cut loop connected to the fourth battery.

34. The electrical system of claim 33, wherein the first and second low voltage cut loops are bundled in a first bundle and the third and fourth low voltage cut loops are bundled in a second bundle; and wherein the first bundle is spaced apart from the second bundle.

35. An aircraft, comprising: a fuselage; a wing connected to the fuselage; a cabin area section of the fuselage; high voltage wiring; battery packs installed in the wing; wherein walls and floor of the cabin area section are free of the high voltage wiring; andwherein the wing includes the high voltage wiring.

36. The aircraft of claim 35, wherein the high voltage wiring runs between the battery packs.

37. The aircraft of claim 35 or 36, wherein the high voltage wiring runs laterally between the fore and aft spars of the wing.

38. The aircraft of any of claims 35-37, wherein the high voltage wiring runs through ribs of the wing.

39. The aircraft of any of claims 35-38, wherein a ceiling of the cabin area section is free of high voltage wiring.

40. The aircraft of any of claims 35-38, wherein a ceiling of the cabin area section that does not include the wing is free of high voltage wiring.

41. The aircraft of any of claims 35-40, wherein a low voltage cut loop extends from at least one of the battery packs through the cabin area section free of high voltage wiring.

42. The aircraft of any of claims 35-41, wherein a tail section of the aircraft is free of high voltage wiring.

43. The aircraft of claim 42, wherein a low voltage cut loop extends from at least one of the battery packs through the tail section of the aircraft free of high voltage wiring.

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