Heat exchanger assembly and cooling system for EVTOL aircraft.

Abstract

The cooling system includes a plurality of heat exchanger assemblies corresponding to a plurality of battery packs (120) and a fluid delivery assembly (1601, 2001, 2101). Each heat exchanger assembly (2200, 2220) includes a first heat exchanger inlet and a second heat exchanger inlet configured to receive or discharge a heat transfer fluid. The fluid delivery assembly (1601, 2001, 2101) is coupled to the heat exchanger assembly and configured to circulate the heat transfer fluid in parallel with the heat exchanger assembly in a U-flow manner, with the heat transfer fluid inlet and outlet being co-located. The fluid delivery assembly (1601, 2001, 2101) includes a plurality of flow restrictors (1630A, 1730A, 1730, 2030) configured to balance the heat transfer fluid flowing into the heat exchanger assembly. The heat transfer fluid flows through a corresponding flow restrictor (1630) before entering a corresponding heat exchanger assembly (2200, 2220) of the battery pack (120).

Description

[Technical Field]

[0001] CROSS-REFERENCE TO RELATED APPLICATIONS This disclosure claims priority to and benefit of U.S. patent application Ser. No. 18 / 460,463, filed Sep. 1, 2023, entitled "HEAT EXCHANGER ASSEMBLIES AND COOLING SYSTEMS FOR EVTOL AIRCRAFT" (Attorney No. 16163.0042-00000), and also claims priority to and benefit of U.S. provisional patent application Ser. No. 63 / 383,660, filed Nov. 14, 2022, entitled "SYSTEMS AND METHODS FOR IMPROVED BATTERY ASSEMBLIES FOR EVTOL AIRCRAFT" (Attorney No. 16163.6005-00000), the contents of which are incorporated herein by reference in their entirety for all purposes.

[0002] Technical Field The present disclosure relates generally to the field of powered air vehicles. More specifically, but not by way of limitation, the present disclosure relates to innovations in aircraft using electric propulsion systems. Certain aspects of the present disclosure generally relate to high-voltage power supply (HVPS) systems and battery assemblies used in air vehicles. Other aspects of the present disclosure generally relate to improvements in cooling distribution systems that may be used in other types of vehicles, but that may provide particular benefits to air vehicles. Summary of the Invention

[0003] According to some embodiments of the present disclosure, a cooling system is provided. The cooling system includes a plurality of heat exchanger assemblies corresponding to a plurality of battery packs and a fluid delivery assembly. Each heat exchanger assembly includes a first heat exchanger inlet and a second heat exchanger inlet configured to receive or discharge a heat transfer fluid. The fluid delivery assembly is coupled to the heat exchanger assemblies and configured to circulate the heat transfer fluid in parallel with the heat exchanger assemblies in a U-flow manner, with the heat transfer fluid inlet and outlet located on the same side. The fluid delivery assembly includes a plurality of flow restrictors configured to balance the heat transfer fluid flowing into the heat exchanger assemblies. The heat transfer fluid flows through the corresponding flow restrictors before entering the corresponding heat exchanger assemblies of the battery packs.

[0004] According to some embodiments of the present disclosure, a power supply system for an aircraft is provided. The power supply system includes: a plurality of battery packs installed in a wing of the aircraft, each battery pack connected to two electric engines of the aircraft to provide power; a plurality of heat exchanger assemblies connected to battery pack enclosures of the plurality of battery packs, respectively; and a fluid delivery assembly coupled to the plurality of heat exchanger assemblies and configured to circulate a heat transfer fluid through the heat exchanger assemblies. Each heat exchanger assembly includes first and second heat exchanger inlets configured to receive and discharge the heat transfer fluid, and a heat exchanger plate including a plurality of triangular dimples configured to provide turbulent flow of the heat transfer fluid.

[0005] According to some embodiments of the present disclosure, a method for cooling a battery is provided. The method includes circulating a heat transfer fluid through a plurality of heat exchanger assemblies corresponding to a plurality of battery packs included in an electric aircraft by a fluid delivery assembly, each heat exchanger assembly including a first heat exchanger inlet and a second heat exchanger inlet configured to receive or discharge the heat transfer fluid, and a heat exchanger plate, and providing turbulent flow of the heat transfer fluid through cooling channels of the heat exchanger plate by a plurality of triangular dimples. The operation of circulating the heat transfer fluid includes, during a first operating period, receiving the heat transfer fluid from the first heat exchanger inlet and discharging the heat transfer fluid to the second heat exchanger inlet, and during a second operating period, receiving the heat transfer fluid from the second heat exchanger inlet and discharging the heat transfer fluid to the first heat exchanger inlet.

[0006] It is to be understood that both the foregoing general description and the following detailed description are exemplary and explanatory only and are not restrictive of the disclosed embodiments as claimed. [Brief explanation of the drawings]

[0007] [Figure 1A] FIG. 1 illustrates a tilt rotor aircraft consistent with certain embodiments of the present disclosure.

[0008] [Figure 1B] FIG. 1 illustrates a tilt rotor aircraft consistent with certain embodiments of the present disclosure.

[0009] [Figure 1C] 1C illustrates an exemplary embodiment of one of the batteries of FIG. 1B.

[0010] [Figure 2] FIG. 1 is a block diagram illustrating a battery pack consistent with some embodiments of the present disclosure.

[0011] [Figure 3]FIG. 1 is a block diagram illustrating aspects of an HV junction box (HVJB) consistent with some embodiments of the present disclosure.

[0012] [Figure 4] FIG. 1 is a cross-sectional view of an aircraft wing having a battery pack installed therein consistent with certain embodiments of the present disclosure.

[0013] [Figure 5A] FIG. 1 illustrates a schematic diagram of a battery pack consistent with some embodiments of the present disclosure.

[0014] [Figure 5B] 1 illustrates components and units within a battery pack consistent with some embodiments of the present disclosure. FIG.

[0015] [Figure 6] FIG. 1 illustrates an exemplary HVJB consistent with certain embodiments of the present disclosure.

[0016] [Figure 7A] 1 illustrates an example circuit diagram of an HVJB consistent with certain embodiments of the present disclosure.

[0017] [Figure 7B] FIG. 1 is an exemplary block diagram illustrating activation of a pyrofuse consistent with some embodiments of the present disclosure.

[0018] [Figure 8] 1A-1C illustrate an example architecture of a foam cell holder consistent with certain embodiments of the present disclosure.

[0019] [Figure 9] FIG. 10 is a top view of a foam cell holder consistent with some embodiments of the present disclosure.

[0020] [Figure 10A]FIG. 1 illustrates an exemplary architecture of a current collecting assembly consistent with certain embodiments of the present disclosure.

[0021] [Figure 10B] 10B shows a top view of the current collecting assembly of FIG. 10A at both the pack level and the cell block level, consistent with some embodiments of the present disclosure. FIG.

[0022] [Figure 11] FIG. 1C is an exemplary circuit diagram illustrating the series and parallel arrangement of battery cells that form the battery pack architecture shown in FIGS. 10A and 10B.

[0023] [Figure 12] 1A-1C illustrate an exemplary vent flap assembly consistent with certain embodiments of the present disclosure.

[0024] [Figure 13A] 13 illustrates an enlarged portion of the vent flap assembly of FIG. 12, consistent with certain embodiments of the present disclosure.

[0025] [Figure 13B] FIG. 13B illustrates a cross-sectional view along line AA of FIG. 13A consistent with some embodiments of the present disclosure.

[0026] [Figure 14] FIG. 10 illustrates a cross-sectional view of a battery pack enclosure including a headspace directly above a vent flap assembly, consistent with some embodiments of the present disclosure.

[0027] [Figure 15] 1 illustrates an exemplary heat exchanger plate consistent with certain embodiments of the present disclosure.

[0028] [Figure 16A] 1 illustrates a design for battery cooling in an HV cooling distribution subsystem consistent with certain embodiments of the present disclosure. [Figure 16B] 1 illustrates a design for battery cooling in an HV cooling distribution subsystem consistent with certain embodiments of the present disclosure. [Figure 16C] 1 illustrates a design for battery cooling in an HV cooling distribution subsystem consistent with certain embodiments of the present disclosure. [Figure 16D] 1 illustrates a design for battery cooling in an HV cooling distribution subsystem consistent with certain embodiments of the present disclosure.

[0029] [Figure 17] 16D illustrates an exemplary battery cooling system based on the refrigerant line architecture of FIG. 16C, consistent with certain embodiments of the present disclosure.

[0030] [Figure 18] 1 illustrates an exemplary home system consistent with certain embodiments of the present disclosure.

[0031] [Figure 19] 19 illustrates an example electronic switching system for the example home system shown in FIG. 18, consistent with some embodiments of the present disclosure.

[0032] [Figure 20] FIG. 1 is a schematic diagram of a cooling system for battery cooling in an HV cooling distribution subsystem consistent with some embodiments of the present disclosure.

[0033] [Figure 21A] 1 illustrates an exemplary design of a cooling system consistent with certain embodiments of the present disclosure. [Figure 21B] 1 illustrates an exemplary design of a cooling system consistent with certain embodiments of the present disclosure. [Figure 21C] 1 illustrates an exemplary design of a cooling system consistent with certain embodiments of the present disclosure. [Figure 21D] 1 illustrates an exemplary design of a cooling system consistent with certain embodiments of the present disclosure. [Figure 21E]1 illustrates an exemplary design of a cooling system consistent with certain embodiments of the present disclosure.

[0034] [Figure 22A] 1 illustrates an exemplary heat exchanger assembly for battery cooling in an HV cooling distribution subsystem, consistent with certain embodiments of the present disclosure.

[0035] [Figure 22B] FIG. 22B is a top view of the heat exchanger assembly shown in FIG. 22A, consistent with certain embodiments of the present disclosure.

[0036] [Figure 22C] 1 illustrates an exemplary refrigerant flow path within a heat exchanger assembly consistent with certain embodiments of the present disclosure.

[0037] [Figure 23A] FIG. 10 illustrates an enlarged portion of a cooling channel in a heat exchanger assembly, consistent with certain embodiments of the present disclosure.

[0038] [Figure 23B] FIG. 23B illustrates a cross-sectional view along line AA of FIG. 23A, consistent with certain embodiments of the present disclosure.

[0039] [Figure 24A] 1A-1C illustrate example dimple placements consistent with certain embodiments of the present disclosure. [Figure 24B] 1A-1C illustrate example dimple placements consistent with certain embodiments of the present disclosure. DETAILED DESCRIPTION OF THE INVENTION

[0040] The following embodiments primarily describe tiltrotor aircraft components used in non-traditional aircraft. For example, tiltrotor aircraft may be intended for frequent (e.g., more than 50 flights per business day) and short-duration (e.g., less than 100 miles per flight) flights over, into, and out of densely populated areas. The aircraft may be intended to carry four to six passengers or commuters who expect a low-noise and low-vibration experience. Therefore, it may be desirable for aircraft components to be configured and designed to withstand frequent use without wear, generate low heat and vibration, and for the aircraft to have mechanisms for effectively controlling and managing heat or vibration generated by the components. Furthermore, several of these aircraft may be intended to operate in close proximity to each other over congested metropolitan areas. Therefore, it may be desirable for aircraft components to be configured and designed to generate low noise levels both inside and outside the aircraft and to have various safety and backup mechanisms.

[0041] For example, for safety reasons, it may be desirable for an aircraft to be propelled by a distributed propulsion system to avoid the risk of single points of failure, and for the aircraft to be able to take off and land normally on a runway. Furthermore, it may be desirable for the aircraft to be able to safely take off and land vertically from and into spaces that are relatively limited compared to traditional airport runways (e.g., vertical airstrips, parking lots, or roadways) while carrying approximately four to six passengers or commuters with their baggage. These usage requirements may impose design constraints on the size, weight, and operational efficiency (e.g., drag, energy consumption) of the aircraft, which may affect the design and configuration of the aircraft's components.

[0042] The disclosed embodiments provide new and improved configurations of aircraft parts not found on conventional aircraft and / or specific design criteria for the parts that differ from conventional aircraft parts. Such alternative configurations and design criteria, combined with addressing shortcomings and challenges of conventional parts, have resulted in the disclosed embodiments of various configurations and designs of tiltrotor aircraft parts.

[0043] In some embodiments, a tiltrotor aircraft may be designed for both vertical and conventional takeoff and landing, while a distributed electric propulsion system enables vertical flight, forward flight, and transition. Thrust is generated by supplying high-voltage power to the electric engines of the distributed electric propulsion system, and each electric engine may convert the high-voltage power into mechanical shaft power to rotate a propeller. Embodiments disclosed herein may include optimizing the energy density of the electric propulsion system. Embodiments may include an electric engine connected to an onboard power source, which may include a device capable of storing energy, such as a battery or capacitor, or one or more systems for harnessing or generating electricity, such as a fuel-driven generator or a solar panel array. Some disclosed embodiments provide for reduced weight and space savings for components within the aircraft, thereby improving the aircraft's efficiency and performance. With a focus on passenger transport safety, the disclosed embodiments implement new and improved safety protocols and system redundancy in the event of a failure, minimizing any single point of failure in the aircraft's propulsion system. Some disclosed embodiments also provide new and improved approaches to meeting aviation and transportation laws and regulations. For example, the Federal Aviation Administration enforces federal laws and regulations requiring safety features, such as firewalls, adjacent to engines that use oil or other flammable materials above a threshold amount.

[0044] In some embodiments, the aircraft may be equipped with a rear engine or lifter that may be of the clockwise (CW) or counterclockwise (CCW) type. Additional embodiments may be equipped with a rear electric engine that uses a multi-blade fixed pitch propeller.

[0045] As described herein, the orientation and use of the electric propulsion systems may vary throughout the operation of the aircraft. In some embodiments, during vertical takeoff and landing, the forward propulsion system and the aft propulsion system may provide vertical thrust during takeoff and landing. During flight phases in which the aircraft is in forward flight mode, the propellers of the aft propulsion system may be retracted in a fixed position to minimize drag, while the forward propulsion system may provide horizontal thrust. The aft electric propulsion system may be actively retracted under position monitoring. Some embodiments may include transitions from vertical flight to horizontal flight and vice versa. In some embodiments, the transition may be performed via a tilt propeller system (TPS). The TPS redirects thrust between a predominantly vertical orientation during vertical flight mode and a mostly horizontal orientation during forward flight mode. Additional embodiments may include a variable pitch mechanism that may change the collective angle of the propeller hub assembly blades of the forward propulsion system for operation during hover, cruise, and transition phases. Some embodiments may include a conventional takeoff and landing (CTOL) configuration, such that the tilter provides horizontal thrust for wing-based takeoff, cruise, and landing. The rear electronic engine is not used to generate thrust during CTOL missions, and the rear propeller is retracted into place.

[0046] In some embodiments, the electric engines described herein may have design features to mitigate and prevent uncontrolled fires, such as utilizing less than one quart or another non-hazardous amount of flammable fluid contained in both the tilt engine and the lift engine, having no nominal ignition source within the electric engine, having an engine overheat operating limit that is more than 50° C. below the auto-ignition temperature of the flammable fluid, overheat detection and protection, overvoltage detection and protection, and overcurrent detection and protection, etc. In some embodiments, the design features of an electric engine may result in it not being in a designated fire zone.

[0047] As disclosed herein, an electric engine may include an inverter and a motor, or an inverter, a gearbox, and a motor in various configurations, such as the exemplary configurations described herein. For example, an electric engine may include an electric motor, a gearbox, and an inverter that all share the same central axis. Additionally, the central axis may be configured along the axis of an output shaft that leads to an aircraft propeller. In such an exemplary configuration, the motor, gearbox, and inverter all share the output shaft as a central axis and are oriented in a circle around the output shaft. Additional embodiments may include a motor, gearbox, and inverter that are mounted together in sequence, or a configuration in which some components, such as the motor and gearbox, are mounted together and other components, such as the inverter, are located elsewhere, and a wiring system is used to connect the electric engine.

[0048] As previously mentioned, the electric engines for aircraft described herein may include some or all of a motor, an inverter, and a gearbox. Various configurations may include an inverter and a motor, such that the motor's output shaft directly provides the speed and torque of the propeller shaft. Additional embodiments of the electric engine may include a motor, inverter, and gearbox, where the motor's output may be transmitted through a gearbox connected to the propeller's output shaft, and a motor, inverter, and gearbox, where the output from the motor is transmitted away from the propeller through the gearbox, and the propeller's output shaft returns to the propeller through the gearbox and motor. As described herein, the electric engine may employ any combination or orientation of some or all of the motor, inverter, and gearbox. Additionally, each configuration or orientation of the electric engine disclosed herein may include cooling via air cooling, a coolant, or a mixture of both.

[0049] For example, an electric engine configuration may include a motor and an inverter, with the motor located between the aircraft's propeller and the inverter. Additionally, the motor may include a gearbox. Furthermore, the inverter may share the same central axis as the motor, and the inverter may be located in a housing cantilevered away from the back of the motor and may be air-cooled. It is recognized that such an inverter orientation results in a less than optimal configuration in terms of the housing required to achieve such a cantilever orientation. Additionally, a motor in this configuration utilizing air cooling may include potting material and air fins to aid in cooling the motor, which may further increase the mass of the system.

[0050] Some embodiments may include an electric engine and the inverter module may be mounted outside the motor housing. Additional embodiments may include an electric engine and the inverter may be mounted on top of the electric motor with the inverter's air cooling fins below the propeller. Further embodiments may include the inverter mounted on the back of the motor with the air cooling fins facing radially outward, the inverter mounted on the front of the motor with the air cooling fins facing radially outward, the inverter mounted on the motor where the inverter is cooled by a liquid such as oil, or other locations of the inverter relative to the motor.

[0051] Embodiments of the electric motor may include a stator housing, a wound stator assembly, a rotor, various bearings, and any additional components to assist in transferring the speed and torque generated by the motor to the propeller.

[0052] It is understood that electric engines can generate heat during operation and may include thermal management systems to ensure that components of the electric engine do not fail during operation. In some embodiments, a coolant may be used and circulated throughout individual components of the engine, such as the inverter, gearbox, or motor, through some of the components, or through all of the engine's components, to help manage the heat present within the engine. Additional embodiments may include using air-cooling methods to cool the electric engine, or a mixture of coolant and air to manage the heat generated during operation of the electric engine. In some embodiments, the coolant used may also be the same liquid used as a lubricant throughout the inverter, gearbox, or motor. For example, the inverter, gearbox, and motor may be cooled using liquid or air, or a mixture of air and liquid cooling may be used, such as using air cooling and liquid cooling in the inverter and gearbox, or any other combination of air and liquid cooling throughout the inverter, gearbox, and motor, or even a subset of these components.

[0053] In some embodiments, oil may be used as a lubricant throughout the electric engine and may also be used as a coolant to help manage heat generated by the engine during operation. Further to this example, different amounts of oil may be used to function as both a lubricant and a coolant within the electric engine, such as less than one quart, less than two quarts, or other amounts of oil required to lubricate and cool the electric engine, with or without the assistance of air cooling. As disclosed herein, electric engines may have different primary functions, such as being used only for takeoff and landing and therefore in only one direction, or being used during all phases of flight, such as takeoff, landing, and flight. An engine used during all phases of flight may experience different orientations throughout flight and may contain more lubricant and coolant than an engine used in only one direction. Therefore, all engines on an aircraft may not contain the same amount of lubricant and coolant. For example, a takeoff and landing engine may require less than one quart of oil, while an engine operating during all phases of flight may require more than one quart of oil. It should be understood that the example embodiments referred to herein are representative and do not define limits on the amounts of lubricants and refrigerants that may be used in an electric engine.

[0054] It is understood that using oil, instead of a separate refrigerant, to not only lubricate but also cool an electric engine adds additional oil to the system, but that oil eliminates traditional components that may be used to cool such an electric engine. For example, if the electric engine is cooled with another liquid, such as glycol, the engine may include separate heat exchangers for both the lubricant and the coolant. Thus, in embodiments where a single fluid, such as oil, is used for both lubrication and cooling, the amount of oil increases, but only one heat exchanger is required. Thus, because fewer heat exchangers are used and other components may be eliminated, the overall system mass may be reduced and a more attractive drag profile may exist. Furthermore, using one substance to lubricate and cool the engine may improve system efficiency due to the reduced mass and the benefits of cooling the engine with a substance instead of relying on air cooling, which can be problematic across the engine.

[0055] Additional embodiments of the electric engine may include various components to monitor and ensure any flammable fluids are prevented from entering certain portions of the electric engine. Some embodiments may include an electric engine with a wet zone enclosure, which may be defined by the gearbox, motor, and / or heat exchanger. In some embodiments, the electric engine may have up to 4 liters of air within the motor gearbox housing that contacts the engine oil. Embodiments of the motor gearbox housing may use a breather to equalize internal and external pressure. Embodiments of the breather may include a breather that protrudes above nearby design features to prevent the inadvertent ingress of external fluids. Additional embodiments may include a breather with a screen and circuitous inlet path to prevent the ingress of external debris. Embodiments may include sight glasses present on both the tilt electric engine and the lift electric engine to check that the oil is not overfilled or underfilled during maintenance.

[0056] Additional embodiments of the electric engine may include active protection features in the front and rear electric engines, such as monitoring internal temperatures during engine operation, including oil temperature, the stator winding set, the inverter bulk capacitor, the power module, the control panel power module, the control panel control processor, the control panel monitor processor, internal hot spots, and various other locations throughout the engine. Embodiments may include overheat limits that account for known fault temperatures and operational limits related to fluid autoignition temperatures. Some embodiments may include a high-voltage power system that may have fuses at the high-voltage battery terminals that may quickly and irreversibly disconnect the engine's electrical connections to mitigate overcurrent events. This overcurrent protection may operate if the current draw of the electric engine is greater than the overcurrent operation. Thus, in some embodiments, a fault condition that leads to an overcurrent may only result in a temporary overheat, arc, or spark fault. Some embodiments may include a fire threat characterization ignition source that may be selected as an ignition source more severe than a short circuit that occurs in the electric engine and is blown by an engine fuse. In some embodiments, the inverter detects AC overcurrent and isolates the error phase and / or continues to monitor the input DC voltage and applies protective action to keep the voltage below the overvoltage operating limit.

[0057] For purposes of this disclosure, the term "substantially" should be understood to mean to a large or significant extent, such as to fall within typical design, machining, and / or manufacturing tolerances understood by one of ordinary skill in the art.

[0058] Reference is made to FIG. 1A , which illustrates a tiltrotor aircraft 100 consistent with some embodiments of the present disclosure. As shown in FIG. 1A , in some embodiments, the distributed electric propulsion system of the tiltrotor aircraft 100 may include twelve electric engines 110, which may be mounted on forward and aft booms of the aircraft 100's wings. The forward electric engines 110 may be tiltable in flight between a horizontally oriented position (e.g., to generate forward thrust) and a vertically oriented position (e.g., to generate vertical lift). The forward electric engines 110 may be clockwise or counterclockwise with respect to the direction of propeller rotation. The aft electric engines 110 may be fixed in a vertically oriented position (e.g., to generate vertical lift) and may also be clockwise or counterclockwise with respect to the direction of propeller rotation.

[0059] Aircraft 100 may have various combinations of forward and aft electric engines 110. For example, in some embodiments, aircraft 100 may have six forward electric engines 110 and six aft electric engines 110. In some other embodiments, aircraft 100 may have four forward electric engines 110 and four aft electric engines 110, or any other combination of forward and aft engines 110. In other embodiments, the number of forward and aft electric engines is not equal.

[0060] In some embodiments, for vertical take-off and landing (VTOL) missions, the forward electric engine 110 and the aft electric engine 110 may provide vertical thrust during take-off and landing. During flight phases in which the aircraft 100 is in forward flight mode, the forward electric engine 110 may provide horizontal thrust, while the propellers of the aft electric engine 110 may be retracted in a fixed position to minimize drag. The aft electric engine 110 may be actively retracted under position monitoring. Transition from vertical to horizontal flight, and vice versa, may be accomplished via a tilt propeller subsystem. The tilt propeller subsystem may redirect thrust between a predominantly vertical orientation during vertical flight mode and a mostly horizontal orientation during forward flight mode. A variable pitch mechanism may vary the collective angle of the propeller hub assembly blades of the forward electric engine for operation during hover, transition, and cruise phases.

[0061] In some embodiments, during conventional take-off and landing (CTOL) missions, the forward electric engine 110 may provide horizontal thrust for wing take-off, cruise, and landing. In some embodiments, the aft electric engine 110 may not be used to generate thrust during CTOL missions, and the aft propeller may be retracted into place.

[0062] In some embodiments, the electric engine 110 is housed in or connected to the boom of the aircraft 100 and may include a motor, an inverter, and a gearbox. In some embodiments, the motor, inverter, and gearbox may be interfaced to share a central axis. In some embodiments, torque generated by the motor may be sent to the gearbox away from the propeller of the propulsion system. In some embodiments, the gearbox may provide gear reduction and then send torque back to the propeller via the main shaft through bearings located inside the motor. In some embodiments, the inverter may be mounted to the rear of the gearbox so that the main shaft does not pass through the inverter when outputting torque to the propeller. In some embodiments, the motor, gearbox, and inverter may be interfaced such that a coolant, such as oil, may be used to service the motor, inverter, and / or gearbox while sharing a common heat exchanger. In some embodiments, the amount of oil used to lubricate and cool the electric engine may vary, including less than 1 quart, less than 2 quarts, less than 3 quarts, or any other measured amount of oil.

[0063] In some embodiments, the tilt propeller system may include a linear or rotary actuator for changing the orientation of the propulsion system during operation. In some embodiments, the pitch of the propulsion system may change depending on the orientation of the propulsion system. In some embodiments, the rotary actuator may include a motor, an inverter, and a gearbox. In some embodiments, the gearbox may include various types of gears interfaced to provide gear reductions that allow the propulsion system to be oriented. In some embodiments, the tilt propeller system may include a redundant configuration in which multiple motors, inverters, and gearboxes are present and interfaced using gears. In some embodiments, a configuration utilizing multiple motors, gearboxes, and inverters may allow a failed portion of the redundant configuration to be driven by a motor, inverter, and gearbox of another portion of the configuration. In some embodiments, the gearbox configuration may also allow the tilt propeller system to maintain the orientation of the propulsion system despite additional power assistance provided by the system.

[0064] As shown in FIG. 1A, the aircraft 100 is configured with a distributed electric propulsion system that enables vertical flight, forward flight, and transition. The six forward electric engines 110 (numbered 1 through 6 from left to right) enable vertical takeoff and landing, transition flight, and full-wing flight through the tilt of controllable pitch propellers. The six aft electric engines 110 (numbered 7 through 12 from left to right) are equipped with fixed-pitch propellers that operate during vertical takeoff and landing and transition and are retracted to a minimum-drag position during normal flight. The flight controls are an integrated fly-by-wire system featuring envelope protection and structural load limiting functions. The aircraft 100 is equipped with advanced cockpit avionics, flight management systems, and sensors necessary to support intended operation and system functions.

[0065] In some embodiments, the electric propulsion system (EPS) described herein may generate thrust by supplying high-voltage (HV) power to the electric engines 110, which in turn convert the HV power into mechanical shaft power used to rotate the propellers. As previously described, the aircraft 100 described herein may have multiple electric engines 110 boom-mounted on the fore and aft wings. The amount of thrust generated by each electric engine 110 may be controlled by a torque command sent from a flight control system (FCS) to each electric engine 110 via a digital communication interface. Embodiments may include forward electric engines 110 that can change their orientation, or cant. Additional embodiments include forward engines that may be clockwise (CW) or counterclockwise (CCW) types. The forward electric engine propulsion subsystem may consist of a multi-blade adjustable-pitch propeller as well as a variable-pitch subsystem.

[0066] In some embodiments, aircraft 100 may include a high-voltage power supply (HVPS) system for providing high-voltage (HV) power. The HVPS system is a power source for aircraft 100 and is configured to distribute stored electrical energy to other systems of aircraft 100, including an electric propulsion system (EPS) that converts electrical power into mechanical rotor power to generate thrust, an environmental control system (ECS) that powers equipment necessary to control the interior atmosphere within the cabin, and a low-voltage system (LVS) that powers equipment and line replaceable units (LRUs) that operate on low-voltage power (e.g., 28 VDC). In some embodiments, the HVPS uses an external HV power source to recharge or replenish the energy sources of aircraft 100. Charging the system ensures that sufficient energy is available to power all equipment, including the ECS, EPS, and low-voltage power supply (LVPS), necessary to successfully execute a subsequent flight mission.

[0067] Specifically, the high-voltage power supply system may include various subsystems, including a battery pack subsystem, an HV vent subsystem, an HV cooling distribution subsystem, and an HV charging subsystem. The battery pack subsystem may provide the necessary power and energy required by aircraft equipment and may include a monitoring unit that measures performance parameters, implements estimation algorithms, and provides active protection against conditions that adversely affect the performance, service life, and safety of the HVPS. The HV vent subsystem is configured to safely remove battery effluent during abnormal operation to prevent structural degradation and minimize the risk of diffusion. The HV cooling distribution subsystem is configured to distribute temperature-conditioned coolant to the battery pack subsystem to achieve the required temperatures for safe charging and discharging operations. The HV charging subsystem may provide an interface between the aircraft 100 and the off-board charger and may be responsible for sequencing charging operations and continuously monitoring voltage and current limits during charging sessions.

[0068] As shown in FIG. 1A , the HVPS system of the aircraft 100 may include multiple, e.g., six, battery packs 120 (numbered 1 through 6 from left to right) installed in battery bays within the wings of the aircraft 100. In some embodiments, to simplify design, manufacturing, and logistics, the battery packs 120 may have identical designs. Each battery pack 120 may be connected to two diagonally opposed electric engines 110 via HV channels 130, for example. HV channel pairs may be determined by their relative positions (left, center, right) on each wing and may be indicated by different dotted / dashed lines in FIG. 1A . For example, the first battery pack 120, numbered 1 in FIG. 1A , may be connected to diagonally opposed electric engines 110, numbered 1 and 12.

[0069] Reference is made to FIG. 1B, which illustrates a tilt-rotor aircraft 195 consistent with some embodiments of the present disclosure. As shown in FIG. 1B, in some embodiments, the tilt-rotor aircraft may comprise a battery assembly including three (3) electrically independent pairs of battery pack units 160, 162, and 164. In some embodiments, each pair of battery pack units 160, 162, 164 may include an electrically independent high-voltage bus 130, 132, 134 connecting pairs of batteries outside the set of six total batteries 101-106 of a given pair of battery pack units 160, 162, 164. FIG. 1C illustrates an exemplary embodiment of one of the batteries 101-106. In some embodiments, each set of paired batteries may power four symmetrical electric engines 170-173, 174-177, or 178-181, which in turn may simultaneously power four electric propeller units (EPUs) (labeled 1-12 in FIG. 1B). In some embodiments, electric engines 170-181 may correspond to the 12 electric engines (labeled 1-12 in FIG. 1A) shown in FIG. 1A.

[0070] 1C, some embodiments may include a pyrotechnic fuse 190 disposed within each one of the batteries 101-106. In some embodiments, the pyrotechnic fuse may be utilized to prevent a single point of failure in the event of a short circuit or failure of any one battery in a set of paired batteries by disconnecting the failed battery after operation of the pyrotechnic fuse, which severs the connection between the failed battery and an operating portion of the tiltrotor aircraft 195's electrical system.

[0071] Referring to FIG. 1B , in some embodiments, additional fuses may be located on the HV bus connecting other components of each pair of battery pack units. For example, fuses 110-121 may be located on the HV bus between each battery 101-106 and its respective electric engine 170-181 and EPU (labeled 1-12 in FIG. 1B ). Each of fuses 110-121 may be utilized to prevent a single point of failure in the event of a short circuit or failure of any one component of one of electric engines 170-181 or EPU (labeled 1-12 in FIG. 1B ), for example, by disconnecting the failed component after operation of a respective one of fuses 110-121, which functions to disconnect the connection between the failed component and an operable portion of the tiltrotor aircraft 195's electrical system. As another example, charging fuses 140, 142, 144, 146, 148, and 150 may be located on the HV bus between each battery 101-106 and a charger or charging system. Each of the charging fuses 140, 142, 144, 146, 148, and 150 may be utilized to further prevent a single point of failure in the event of a short circuit or failure of any one battery during charging, for example, by disconnecting the failed battery from the charger or charging system after operation of one or more of the fuses 140, 142, 144, 146, 148, and 150, which severs the connection between the failed battery and an operable portion of the electrical system of the tiltrotor aircraft 195.

[0072] As described herein, the orientation and use of the electric propulsion system may vary throughout the operation of the aircraft in response to various failure scenarios. In some embodiments, each set of paired batteries may simultaneously power four symmetrical electric engines.

[0073] 1B and 1C , the first battery 101 of the first pair of battery pack units 160 may directly power two of the twelve electric engines 170, 171, which in turn may power two EPUs (labeled 1 and 12 in FIG. 1B ), and the second battery 102 (not shown) of the pair of battery pack units 160 may directly power two of the twelve electric engines 172, 173, which in turn may power two of the twelve EPUs (labeled 4 and 9 in FIG. 1B ). In some embodiments, the first battery 101 and the second battery 102 (not shown) of the pair of battery pack units may be cross-linked (i.e., electrically interconnected) via a first high-voltage (HV) cross-link bus 130. In some embodiments, the first HV cross link bus 130 enables the first battery 101 to function as a backup battery for the second battery 102 (not shown), thereby powering the two electric engines 172, 173 and the EPU (labeled as 4 and 9 in FIG. 1B ) initially powered by the second battery 102 when the second battery 102 (not shown) fails. In some embodiments, the first HV cross link bus 130 enables the second battery 102 (not shown) to function as a backup battery for the first battery 101, thereby powering the two electric engines 170, 171 and the EPU (labeled as 1 and 12 in FIG. 1B ) initially powered by the first battery 101 when the first battery 101 fails. In some embodiments, if either the first battery 101 or the second battery 102 (not shown) of the first pair of battery pack units 160 fails, the failed battery may be disconnected from the other working batteries of the pair of battery packs.In some embodiments, disconnection may be performed via a pyrotechnic fuse 190 located near the output of each battery 101-106, and thus along the HV connection between the first battery 101 and the second battery 102 of the paired battery pack units 160. In some embodiments, each battery 101, 102 of the first paired battery pack units 160 may be electrically separable from the rest of the first paired battery pack units 160 and from any other paired battery pack units 162, 164, for example, via charging fuses 140, 142, so that a fault may be an isolated event that does not propagate along any connected HV channel.

[0074] Continuing with the above example and further referring to FIGS. 1B and 1C , the third battery 103 (not shown) of the second pair of battery pack units 162 may directly power two additional electric engines 174, 175, which may in turn power two other EPUs (labeled 2 and 11 in FIG. 1B ), and the fourth battery 104 (not shown) of the second pair of battery pack units 162 may directly power two other electric engines 176, 177, which may in turn power two other EPUs (labeled 5 and 8 in FIG. 1B ). In some embodiments, the third battery 103 and the fourth battery 104 of the second pair of battery pack units 162 may be cross-linked (i.e., electrically interconnected) via a second HV cross-link bus 132. In some embodiments, the second HV cross link bus 132 may enable the third battery 103 to function as a backup battery for the fourth battery 104, thereby powering the two electric engines 176, 177 and the EPU (labeled as 5 and 8 in FIG. 1B ) initially powered by the fourth battery 104 when the fourth battery 104 fails. In some embodiments, the second HV cross link bus 132 may enable the fourth battery 104 to function as a backup battery for the third battery 103, thereby powering the two electric engines 174, 175 and the EPU (labeled as 2 and 11 in FIG. 1B ) initially powered by the third battery 103 when the third battery 103 fails. In some embodiments, if either the third battery 103 or the fourth battery 104 of the second pair of battery pack units 162 fails, the failed battery may be disconnected from the other operational batteries of the second pair of battery pack units 162. In some embodiments, the disconnection may be performed via a pyrotechnic fuse 190 located near the output of each one of the batteries 101-106, and thus along the HV connection between the third battery 103 and the fourth battery 104 of the second pair of battery pack units 162.In some embodiments, for example, each battery 103, 104 of the second pair of battery pack units 162 may be electrically separable from the remainder of the pair of battery pack units 162 and any other pairs of battery pack units 160, 164, e.g., via charging fuses 144, 146, so that a fault may be an isolated event that does not propagate along any connected HV channels.

[0075] 1B and 1C , the fifth battery 105 (not shown) of the third pair of battery pack units 164 may directly power two additional electric engines 178, 179, which in turn may power two other EPUs (labeled 3 and 10 in FIG. 1B ), and the sixth battery 106 of the third pair of battery pack units 164 may directly power two other electric engines 180, 181, which in turn may power two other EPUs (labeled 6 and 7 in FIG. 1B ). In some embodiments, the fifth battery 105 and the sixth battery 106 of the third pair of battery pack units 164 may be cross-linked (i.e., electrically interconnected) via a third HV cross-link bus 134. In some embodiments, the third HV cross link bus 134 may enable the fifth battery 105 to function as a backup battery for the sixth battery 106, thereby powering the two electric engines 180, 181 and the EPU (labeled as 6 and 7 in FIG. 1B ) initially powered by the sixth battery 106 when the sixth battery 106 fails. In some embodiments, the third HV cross link bus 134 may enable the sixth battery 106 to function as a backup battery for the fifth battery 105, thereby powering the two electric engines 178, 179 and the EPU (labeled as 3 and 10 in FIG. 1B ) initially powered by the fifth battery 105 when the fifth battery 105 fails. In some embodiments, if either the fifth battery 105 or the sixth battery 106 of the third pair of battery pack units 164 fails, the failed battery may be disconnected from the other operational batteries of the third pair of battery pack units 164. In some embodiments, the disconnection may be performed via a pyrotechnic fuse 190 located near the output of each one of the batteries 101-106, and thus along the HV connection between the fifth battery 105 and the sixth battery 106 of the third pair of battery pack units 164.In some embodiments, for example, each battery 105, 106 of the third pair of battery pack units 164 may be electrically isolated from the rest of the pair of battery pack units 164 and any other pairs of battery pack units 160, 162, for example via charging fuses 148, 150, so that a fault may be an isolated event that does not propagate along any connected HV channels.

[0076] 1B , in some embodiments, as a result of the above example configuration, three pairs of battery pack units (i.e., first, second, and third pairs of battery pack units) 160, 162, 164 may enable optimization of the power available to operate the 12 electric engines (labeled 1-12 in FIG. 1B ) by configuring each pair of battery pack units 160, 162, 164 to power a specific one of the 12 electric engines. As an example, in some embodiments, each one of the three pairs of battery pack units 160, 162, 164 may be configured such that each set of batteries in a pair may power any four of the 12 electric engines, e.g., one of the forward outer electric engines, one of the aft inner electric engines, one of the forward inner electric engines, and one of the aft outer electric engines. As another example, in some embodiments, each battery pack unit 160, 162, 164 may be configured such that each set of paired batteries may power any combination of four forward and / or aft electric engines and / or EPUs. Appendix 1 provides additional information regarding aspects of the high voltage power supply (HVPS) system.

[0077] In some embodiments, as a result of the above exemplary design, particularly the ability of one battery in each pair of battery packs to serve as a backup battery for a failed battery in that pair of battery packs, the backup battery may simultaneously power four electric engines (i.e., the backup battery may power two electric engines that are electrically connected directly to the operating battery and two electric engines via an HV crosslink bus connecting the paired battery sets of the paired battery pack units). In some embodiments, the backup battery may also simultaneously power the four electric engines at a slightly reduced power level that maintains a power level sufficient for proper flight control. As a result of such an exemplary design, all electric engines continue to receive power, and thus all motors and propellers continue to operate, even if the first battery in any or all of the first, second, and third pair of battery pack units fails. As a further result of such an exemplary design, generally about 20% less battery power may be required compared to an independent bus or common bus architecture to achieve the same results (e.g., to reach sufficient energy to controllably fly the design mission and reach sufficient thrust to safely execute a hover landing). In some embodiments, as a further result of such an exemplary design and an approximately 20% reduction in power requirements, improved energy efficiency may also be achieved, for example, due to the ability to be powered for longer periods of time, extending the aircraft's range, and reducing the weight of the aircraft due to the ability to utilize lower power batteries.

[0078] It is understood that in some embodiments, a battery pack unit may include one battery that powers any two electric engines and EPUs, and an aircraft may include six battery pack units. As another example, in some embodiments, a battery pack unit may include three connected batteries, each of which may directly power any two electric engines and EPUs, and each battery pack unit with three connected batteries may power any six total electric engines and EPUs of the aircraft. As yet another example, in some embodiments, a battery pack unit may include four connected batteries, each of which may directly power any two electric engines and EPUs, and each battery pack unit with four batteries may power any eight total electric engines and EPUs of the aircraft. It is understood that any number of electric engines and EPUs may refer to electric engines and EPUs located on both sides of the aircraft, at the front of the aircraft, and / or at the rear of the aircraft. It is further understood that in some embodiments, any combination of different sized battery pack units may be utilized to power 12 electric engines and EPUs. It will further be appreciated that any combination of different sized battery pack units may be utilized to power any number of electric engines and EPUs on an aircraft, with no limit on the total number of electric engines and EPUs.

[0079] Please refer to Figure 2, which is a block diagram illustrating a battery pack 120 consistent with some embodiments of the present disclosure. As shown in Figure 2, the battery pack 120 may be electrically and communicatively connected to an HV load 210, an HV vent subsystem 220, an FCS 230, a low voltage power supply (LVPS) system 240, an aircraft switch 250, an HV charging subsystem 260, a ground support system (GSS) 270, and an LV power source 280.

[0080] In the embodiment of FIG. 2, the battery packs 120 include an HV junction box (HVJB) 122 and a cell stack assembly 124. Each battery pack 120 includes an HV distribution unit (HVJB) and a battery management system (BMS) housed within the HVJB 122. Specifically, the BMS may be configured to monitor voltage, temperature, current, and insulation resistance and control pack contactors and pyro-fuses to protect against fault conditions for safe operation. The BMS may also execute algorithms to determine the status of the pack (e.g., state of charge, state of health, etc.). The BMS may receive voltage and temperature sensing signals from the cell stack assembly 124 and the HV distribution unit and control the HV distribution unit accordingly. The HV loads 210 receiving HV power may include an electric engine (e.g., two corresponding electric engines 110 connected to the battery packs 120 shown in FIG. 1A), a tilt actuator, a DC / DC converter, an ECS, or a cross-link used to connect the pack pair.

[0081] Battery pack 120 may be connected to FCS 230 via a digital communication bus. Battery pack 120 may receive a low-voltage input voltage (e.g., 28V input) from LVPS system 240. Battery pack 120 may receive digital input signal(s) from aircraft switch 250 and provide a supply voltage (e.g., 28V input) to aircraft switch 250. HV power can be transmitted between GSS 270 and battery pack 120 via HV charging subsystem 260. In some embodiments, HV charging subsystem 260 includes a charge port assembly 262 having a control MCU (CCU) connected to LV power source 280 to provide a supply voltage (e.g., 28V input) to HV charging subsystem 260. In some embodiments, GSS 270 may include a thermal regulation subsystem 272 and a ground charging subsystem 274 electrically and communicatively connected to charge port assembly 262.

[0082] Reference is now made to Figure 3, which is a more detailed block diagram 300 illustrating units within HVJB 122 consistent with certain embodiments of the present disclosure. As shown in Figure 3, in some embodiments, HV distribution unit 310 within HVJB 122 may include HV contactor 312 and a combination of active and passive fuses (e.g., pyrofuse 314 and fuse 316) to protect against overcurrent and short-circuit conditions.

[0083] The BMS 320 in the HVJB 122 may include a cell management unit (CMU) 322 for monitoring the voltage of each set of seven series-connected parallel cells (i.e., 1S-7P cell groups) in the 14S-7P cell block. The CMU may also be used to monitor the temperature of the 14S-7P cell block. The CMU 322 obtains measurements of all cell groups in the battery pack 120 and transmits the measurements to the battery management unit (BMU) 324 via an isolated serial peripheral interface (isoSPI) in a daisy-chain configuration. In some embodiments, the CMU 322 does not have any active management or control mechanism for the cells in the cell block, but the CMU 322 can perform passive cell balancing of the series cell block when instructed by the BMU 324. The BMU system architecture may provide the flexibility to instruct passive balancing both on the ground and in the air.

[0084] The BMU printed circuit board assembly (PCBA) is configured to monitor the voltage, current, and temperature from the CMU 322, in addition to the output current to each of the connected loads. The BMU 324 is internally powered by the battery cell stack assembly 124 and may continue to monitor the condition of the batteries even when not onboard the aircraft 100. By monitoring parameters of the battery pack 120, cell blocks, and cell groups, the BMU can protect against conditions that adversely affect safety or performance, such as overvoltage, undervoltage, overheating, undertemperature, loss of electrical insulation, short circuits, and overcurrent.

[0085] The diagnostic capabilities of the BMU 324 enable fault detection and isolation through built-in test (BIT). Fault indications and overall usage are logged into non-volatile memory (NVM) for data retention and prediction. Additionally, the BMU 324 performs calculations of the battery pack 120's state of charge (SOC), state of health (SOH), state of power (SOP), state of energy (SOE), and state of temperature (SOT). The BMU 324 also controls and monitors bus pre-charging and provides contactor commands. In some embodiments, the BMU 324 may use two microcontrollers to perform state calculations and contactor and fuse control functions so that a subset of BMU 324 functionality is preserved when a single microcontroller fails.

[0086] The CCU 330 in the charging port assembly 262 may interface with the external battery charger and communicate with the BMUs 324 on the six attached battery packs 120. In some embodiments, the CCU 330 may be a single PCBA with one microcontroller that manages the overall power supply to each battery pack 120 during charging. As shown in FIG. 2 , the CCU 330 may perform the handshake between the ground charging subsystem 274 and the BMUs 324, request that the BMUs 324 open or close contactors 312, and provide active detection and protection functions for overvoltage protection. In some embodiments, the CCU 330 is external to the battery packs 120. The BMUs 324 in each battery pack 120 maintain complete control and monitoring of those battery packs 120 during charging operations.

[0087] Reference is made to Figure 4 and Figures 5A and 5B. Figure 4 is a cross-sectional view of a wing 400 of the aircraft 100 of Figure 1A within which a battery pack 120 is installed, consistent with some embodiments of the present disclosure. In some embodiments, the battery pack 120 can be removed from service when it reaches a state of health where it no longer meets the minimum energy and power requirements necessary for operation. As shown in Figure 4, in some embodiments, an elbow duct 410 is connected to a vent 420 to exhaust battery waste upward. The HVJB 122 of the battery pack 120 is attached to the battery pack enclosure of the cell stack assembly 124.

[0088] FIG. 5A is a diagram illustrating a schematic diagram of a battery pack 120 consistent with some embodiments of the present disclosure. FIG. 5B is a diagram illustrating components and units within the battery pack 120 of FIG. 5A consistent with some embodiments of the present disclosure. In FIGS. 5A and 5B, the battery pack 120 is shown upside down relative to the installation orientation shown in FIG. 4. As shown in FIGS. 5A and 5B, the battery pack 120 includes an HVJB 122 and a cell stack assembly 124. The cell stack assembly 124 includes an enclosure 510 including a heat exchanger (e.g., a heat exchanger plate), a cell holder 520 including interstitial foam, a plurality of battery cells 530 (e.g., 1456 cells in each battery pack 120), a current collector assembly 540, a vent flap assembly 550, a pack bus bar 560, and a pack lid 570. As shown in FIG. 5A, the HVJB 122 may be attached to the pack lid 570 using fasteners. Details of the cell stack assembly 124 are described in the following paragraphs with reference to the accompanying drawings.

[0089] In some embodiments, the battery cells 530 in the battery pack 120 may be lithium-ion cells, although the present disclosure is not limited thereto. For example, the battery cells 530 may be 4.5 Ah capacity cells with a 2170 cylindrical form factor. The battery cells 530 can operate over a voltage range of approximately 2.5 V to approximately 4.18 V, and each cell has a beginning-of-life (BOL) capacity of 4.5 Ah. In some embodiments, each battery cell 530 includes a current interrupt device (CID) that permanently places the battery cell 530 in an open circuit state if internal pressure increases.

[0090] 6 illustrates an exemplary HVJB 122 consistent with certain embodiments of the present disclosure. As discussed above, the HVJB 122 is configured to distribute high-voltage stored electrical energy for propulsion and other system needs and may include pre-charge resistors and relays, HV bus bars, shunt resistors, HV contactors, connectors, a combination of active and passive fuses to protect against overcurrent and short-circuit conditions, and components to disable high-voltage output in an emergency if deemed necessary. The components within the HVJB 122 may be encapsulated with a BMU PCBA and attached to the battery pack enclosure using fasteners.

[0091] 7A shows an example circuit diagram 700A of the HVJB 122 consistent with certain embodiments of the present disclosure. As described above, the HVJB 122 may be electrically connected to the HV load 210 to provide HV power. Specifically, the HV power may be provided using a DC / DC converter 710 in the BMS 320 and a power storage element BT1 (e.g., battery cells 530 connected in parallel and series).

[0092] As shown in FIG. 7A , the DC / DC converter 710 and power storage element BT1 are connected to each of the HV loads 210 via pre-charge resistor(s) (e.g., resistor R1) or current-sensing resistor(s) (e.g., resistors R2-R6), switching devices K1-K7 (e.g., HV contactors), and a combination of active and passive fuses (e.g., pyrofuses F2, F3, F4, and F6, and fuses F1, F5, F7, and F8) to protect against various fault conditions (e.g., overcurrent, short circuit, etc.). The active and passive fuses may be located within the HV battery pack 120 to safely cut off power from the HV battery pack 120 in the event of a short circuit on the load side. The HV loads 210 may include two corresponding electric engines 110, a tilt actuator 212, a DC / DC converter 214 (e.g., a 2.8 kW DC / DC converter), and an ECS load 216.

[0093] For example, pyrofuses F2, F3, F4, and F6 may be a type of fuse configured to be operated by an external power source when circuit disconnection and isolation is required. BMS 320 is used to operate pyrofuses F2, F3, F4, and / or F6 when a short-circuit event occurs. In some embodiments, BMS 320 in HVJB 122 can send a command signal to a corresponding pyrofuse driver to operate the pyrofuse when a fault occurs, thereby protecting against overcurrent and electrically isolating the entire battery pack 120 from the connected load. In some embodiments, HVJB 122 may further provide redundant active triggers configured to enable the pyrofuse driver to operate one or more pyrofuses if BMS 320 fails to enable the pyrofuse driver, as described below with reference to the accompanying drawings.

[0094] FIG. 7B is an example block diagram 700B illustrating pyrofuse activation consistent with some embodiments of the present disclosure. As shown in FIG. 7B , in some embodiments, the voltage output from the battery cell stack assembly 124 can be HV (e.g., 520V to 870V). This voltage can be converted by an internal DC-DC converter 710 within the battery pack 120 to power the BMS 320 and a pyrofuse redundant trigger (PRT) board 730 located within the HVJB 122. The PRT board 730 can include various analog components to form a plurality of control logic circuits 732 to provide redundant active triggering for corresponding pyrofuses (e.g., pyrofuses F2, F3, and F6). In some embodiments, the battery cell stack assembly 124 can also be electrically connected to an external DC-DC converter 720 located outside the battery pack 120. The external DC-DC converter 720 can also be configured to convert the voltage output from the battery cell stack assembly 124 to provide supply voltages to the BMS 320 and the PRT board 730. 7B, the power supply may be ensured and the stability of the system may be improved by a redundant power supply system including an internal DC-DC converter 710 and an external DC-DC converter 720. If the main power converter (which may be either the internal DC-DC converter 710 or the external DC-DC converter 720) fails, the system can switch to the backup power converter to power other devices and units in the battery pack 120 to ensure that the system operates normally.

[0095] 7B , multiple pyrofuse drivers 734 configured to operate corresponding pyrofuses (e.g., pyrofuses F2, F3, and F6) may be disposed on PRT board 730, although the disclosure is not limited thereto. In other embodiments, pyrofuse drivers 734 may also be disposed separately on another circuit board. When BMS 320 detects an overcurrent fault based on a signal sensed by a corresponding current-sensing resistor (e.g., resistor R3, resistor R4, or resistor R6), BMS 320 may output a corresponding pyrofuse control command signal (e.g., EE1 PYRO CONTROL for a pyrofuse driver associated with a power line to a first electric engine, EE2 PYRO CONTROL for a pyrofuse driver associated with a power line to a second electric engine, or XLINK PYRO CONTROL for a pyrofuse driver associated with a power line to a crosslink used to connect the pack pair). Thus, after receiving the pyrofuse control command signal, the corresponding pyrofuse driver 734 may activate the corresponding pyrofuse to electrically isolate the battery pack 120 from the load.

[0096] In some embodiments, if the BMS 320 fails to function and output a pyrofuse control command signal when a fault occurs, the pyrofuses will not blow and disconnect the power path. In such cases, the control logic circuit 732, configured with different thresholds for triggering the pyrofuses, may act as a redundant trigger and output a corresponding command signal to the pyrofuse driver 734 to activate the corresponding pyrofuse. Thus, the PRT board 730 is completely independent of the BMS 320 and is designed to trigger and blow the pyrofuses based on analog circuitry. The redundant trigger source can further increase the safety level of the HV battery pack 120.

[0097] 7B, in some embodiments, both the BMS 320 and the control logic circuit 732 within the PRT board 730 may use the same current sensors (e.g., sense resistor R3, resistor R4, and resistor R6) to detect short-circuit conditions. The BMS 320 may be the primary source of sending activation signals to the pyrofuses, and the PRT board 730 may be responsible for confirming safety events and taking action only if the BMS 320 fails to activate the pyrofuses.

[0098] Reference is made to FIG. 8, which illustrates an exemplary architecture of a foam cell holder 520 consistent with some embodiments of the present disclosure. In some embodiments, the cell holder 520 can be machined foam with an FR4 top sheet and embedded bus bars. In some embodiments, the cell holder 520 can be a precision-manufactured, rigid, flame-retardant, closed-cell polyurethane foam structure. Various properties are considered when selecting a material for the foam cell holder 520, including, for example, mechanical rigidity (e.g., elastic modulus, yield strength, etc.), minimal thermal conductivity, ease of adhesive bonding, high dielectric constant, high dielectric breakdown resistance, low density (or low weight), and / or mass production capabilities (e.g., ability to be cast into complex shapes while minimizing waste). For example, various structural foams may meet aerospace industry standards. The cell holder 520 may be resistant to water absorption and mold growth, and may be self-extinguishing and flame-retardant. The closed-cell structure of the cell holder 520 acts as an insulator and gas barrier to minimize conduction and convection between the battery cells 530. Thus, the foam cell holder 520 positions the battery cell 530 and separates the battery cell 530 from adjacent cells to protect the battery cell 530 from thermal runaway conditions. The cell holder 520 may also include nested bus bars for connecting portions of the battery pack 120.

[0099] FIG. 9 is a top view of a foam cell holder 520 consistent with some embodiments of the present disclosure. In some embodiments, the foam cell holder 520 can be configured to accommodate a total of 1,456 cylindrical battery cells 530. For example, the battery cells 530 may be arranged into five columns 910-950. Each of the columns 920-940 includes 42 rows of seven cells, for a total of 294 cells per column. The first column 910 and the last column 950 have one fewer row, each with a total of 287 cells. Each row of seven parallel cells can be referred to as a cell group 960. Each cell is insulated from adjacent cells in the same or adjacent cell groups by a nominal 1.5 mm of foam to reduce conductive heat transfer and the possibility of cell sidewall damage.

[0100] In some embodiments, the foam cell holder 520 includes registration features and holes for positioning the part relative to the manufacturing fixture. Channels along the edges of the foam are used as mounting locations for separate aluminum bus bars that are bolted and welded to the current collecting assembly 540.

[0101] FIG. 10A illustrates an exemplary architecture of a current collecting assembly 540 consistent with some embodiments of the present disclosure. In some embodiments, the current collecting assembly 540 may be a flexible circuit current collecting assembly with integrated sensing components glued and laser welded. In some embodiments, the sensing components may be integrated directly onto a flexible printed circuit board that is part of the current collecting assembly, without the need for glue or welding. As shown in FIG. 10A , the current collecting assembly 540 may include five columns of cells and components integrated onto a single central cell holder 520. In some embodiments, the current collecting assembly 540 may implement a laminated busbar with an integrated sensing layer for sensing the voltage and / or temperature of a corresponding cell group 542 or a corresponding cell block 544, providing a single component that integrates the current collectors and sensing lines through lamination. For example, in each column, voltage and temperature sensing lines 546 may be arranged through lamination and configured to collect voltage traces at one end of the column.

[0102] In some embodiments, the current collecting assembly 540 is bolted and welded to an aluminum bus bar 548, along with a main pack bus bar 560 used to connect the positive and negative ends of the cell stack assembly 124 to the HVJB 122. In some embodiments, the current collecting assembly 540 is fabricated with nickel-plated copper conductors that are attached to the individual battery cells 530 to form electrical paths for the desired cell combinations. For example, the current collecting assembly 540 has a 208S-7P pack architecture with a series of 208 cell groups 542, with each cell group 542 comprising seven parallel cells. The current collecting assembly 540 may integrate both voltage and temperature sensing lines for each 14S-7P cell block 544. Each cell block 544 comprises a series of 14 cell groups 542. That is, there are three cell blocks 544 per column, for a total of 15 cell blocks 544 in the current collecting assembly 540.

[0103] 10B illustrates a top view of the current collecting assembly 540 of FIG. 10A at both the pack level and the cell block level, consistent with some embodiments of the present disclosure. The current collecting assembly 540 may be attached to the foam cell holder 520 using a pressure-sensitive adhesive (PSA). The battery cells 530 within the foam cell holder 520 are welded to the current collecting assembly 540 to form an electrical connection.

[0104] The current collecting assembly 540 may form both parallel and series connections of the battery pack 120, allowing for electrical monitoring of each cell group and thermal monitoring of each cell block. With the 208S-7P pack architecture, if an internal short circuit occurs in a battery cell 530, only battery cells 530 within the same 7P cell group are affected by the short circuit. FIG. 11 is an exemplary circuit diagram illustrating the series and parallel arrangement of battery cells 530 forming the 208S-7P pack architecture shown in FIGS. 10A and 10B. In some embodiments, with this 208S-7P pack architecture shown in FIG. 11, the battery pack 120 is configured to provide a maximum voltage of 874V (i.e., 4.2V for each cell group 542).

[0105] In some embodiments, an insulating vent flap may be placed over the welded current collecting assembly 540. Reference is made to Figures 12 and 13A and 13B. Figure 12 illustrates an exemplary vent flap assembly 550 consistent with certain embodiments of the present disclosure. Figure 13A illustrates an enlarged portion of the vent flap assembly 550 of Figure 12 consistent with certain embodiments of the present disclosure. Figure 13B illustrates a cross-sectional view along line AA of Figure 13A consistent with certain embodiments of the present disclosure. In Figures 13A and 13B, the vent flap assembly 550 is shown upside down relative to the installation orientation.

[0106] Specifically, as shown in FIGS. 12 and 13A and 13B , the vent flap assembly 550 includes two layers 552 and 554 that are flexible and capable of withstanding high temperatures (e.g., temporary or sustained high temperatures). For example, during a thermal runaway event, the temperature may reach approximately 800° C. The first layer 552 may be a flexible silicone foam layer with a PSA used to secure the vent flap assembly 550 to the current collecting assembly 540, although the present disclosure is not limited thereto. The second layer 554 is configured to protect adjacent cell groups 542 from cell blowout and may be formed using flexible ceramic, thermoplastic polyimide, or any other material of the same class that shares similar properties and characteristics. In addition to the star shape shown in FIG. 12 , various designs can be used to implement the vent flap assembly 550. For example, in some other embodiments, the vent flap assembly 550 may use relatively rigid, non-cantilevered elements to provide a structure that may rupture under pressure without significant outward extension during a thermal runaway event. The ventilation flap assembly 550 may adopt various shapes that allow deformation of the cantilevered portion(s) to allow ventilation, for example, a honeycomb structure may be applied to the arrangement of the battery cells, but the present disclosure is not limited thereto.

[0107] 13A and 13B show the series connection in addition to the vent manifold 556 and cell headspace 558 in the vent flap assembly 550. In some embodiments, the cell headspace 558 provides an open volume (e.g., approximately 20 mm in height) between the pack lid 570 and the second layer 554 to vent gases and reduce convective heat transfer between different cell groups. In some embodiments, the battery pack 120 may also provide a battery pack vent (e.g., another one-way valve), which may be a commercially available device. In some embodiments, when a cell releases gas, the released gas from the affected cell enters the space below the cell and is contained and directed downward by the foam cell holder 520, which is surrounded by the silicone foam layer 552 and the polyamide thermoplastic layer 554. This space may be referred to as the vent manifold 556 and is interconnected for cells belonging to one cell group. The vent manifold 556 is not interconnected to other cell groups because it is separated by the PSA-compatible silicone foam layer 552. As pressure increases within the vent manifold 556, a flap on the polyamide thermoplastic layer 554 opens, allowing gas and potential ejecta to enter the headspace 558 below the vent flap assembly 550. The vent flap assembly 550 may thus provide a one-way valve located above the cell and current collector assembly 540 to safely vent a cell experiencing thermal runaway while protecting adjacent cells from the heat and ejecta of the failed cell.

[0108] FIG. 14 shows a cross-sectional view of a battery pack enclosure 1400 including a headspace directly above the vent flap assembly 550, consistent with some embodiments of the present disclosure. Similar to FIGS. 13A and 13B, the structure in FIG. 14 is shown upside down relative to the installation orientation. As shown in FIG. 14, the battery pack enclosure 1400 provides a structural interface for attaching the battery pack 120 to the primary structure of the aircraft, houses the entire 208S-7P cell stack assembly 124, and connects to a heat exchanger. In some embodiments, the battery pack enclosure 1400 may provide a 20 mm headspace 558 between the pack lid 570 and the cell stack to reduce convective heat transfer between cell groups.

[0109] Thus, the battery pack enclosure 1400 includes a defined path for venting vent gases via a pressure relief safety device (e.g., a burst valve) that prevents over-pressurization of the battery pack 120. The pack vent may be a two-stage system used to provide pressure equalization between the battery pack 120 and the ambient environment during normal operation. The vent interfaces with a duct that exhausts battery exhaust to the exterior of the aircraft 100. Aside from the vent path, the battery pack enclosure 1400 may be sealed to ensure that any cell electrolyte leaks are contained within the enclosure and do not flow into the aircraft 100 where they could pose a hazard.

[0110] FIG. 15 illustrates an exemplary heat exchanger plate 1500 consistent with some embodiments of the present disclosure. The heat exchanger plate 1500 may be secured to the battery pack enclosure 510 to enable regulation of the temperature within the battery pack 120 using a thermally conditioned fluid for heating and cooling, such as ethylene glycol water (EGW), ethylene glycol, polyethylene glycol, water, or any combination thereof. In some embodiments, active thermal regulation may be used only while the aircraft 100 is on the ground. The heat exchanger plate 1500 may be sized to support the required heating or cooling needs while operating under nominal pressure and flow conditions. For example, the heat exchanger plate 1500 may be bonded to the battery pack enclosure 510 using a thermally conductive adhesive (e.g., a flame-retardant acrylic adhesive). The thermally conductive adhesive may be used to conduct heat generated by the cells to the heat exchanger plate 1500.

[0111] 16A-16D each illustrate different designs for battery cooling in an HV cooling distribution subsystem consistent with some embodiments of the present disclosure. Some item numbers are shown in FIG. 16A and may be omitted in subsequent figures for clarity. As previously discussed, the aircraft 100 may include, for example, six battery packs 120. In some embodiments, a different number of battery packs 120 may be provided, such as two, four, eight, or ten battery packs 120. In some embodiments, an even number of battery packs may be provided to achieve symmetrical weight distribution within the aircraft 100. In some embodiments, an odd number of battery packs 120 may be provided. For example, for a configuration in which the battery packs 120 can be safely accommodated in a central location of the aircraft 100, a central battery pack 120 may then be provided with an equal number of battery packs 120 arranged on either side of the central battery pack 120.

[0112] Each battery pack 120 may include a heat exchanger inlet 1621 and a heat exchanger outlet 1622 for circulating refrigerant to provide battery cooling. Figures 16A-16D illustrate different refrigerant line architectures that may provide a single main inlet 1610 and main outlet 1650, respectively, within the aircraft 100. The main inlet 1610 and main outlet 1650 may be configured to receive refrigerant from a ground-based refrigerant supply system and discharge the refrigerant thereto, circulating the refrigerant via the inlet path 1611 and the outlet path 1651. The inlet path 1611 and the outlet path 1651 may comprise, for example, rigid or semi-rigid piping, flexible hoses, integrally molded channels, or conduits formed from a combination of these or other suitable materials. The collection of such piping and associated fixtures, fittings, inlets, and outlets for distributing refrigerant to the heat exchangers within each battery pack 120 may be referred to as a fluid transfer assembly. Configurations according to embodiments of the present disclosure enable even and parallel distribution of coolant to multiple battery packs 120 installed within the wings of an aircraft 100 without the need for active control components. The arrangements of FIGS. 16A-16D may be configured to reduce the overall diameter, length, and number of fittings of the inlet and outlet paths 1611 and 1651, thereby minimizing both the dry and wet weight of the cooling system. Providing a single main inlet and outlet within the aircraft 100 allows for quick connection and disconnection to and from a ground-based coolant supply, thereby enabling rapid turnaround between landing and takeoff. This may advantageously reduce downtime in flight schedules. Alternatively, it may effectively extend cooling time for a given landing period. In some embodiments, battery cooling may be achieved without the use of onboard pumps or other active flow control components. Instead, an external pump (e.g., on a ground-based coolant supply system external to the aircraft 100) may be used to circulate coolant within the battery packs 120, for example, while simultaneously charging the battery packs 120.

[0113] In some embodiments, the cooling operation may occur without exceeding the full system pressure limits (e.g., less than 100 psi), thus eliminating the need for special connectors. The cooling system may be a passive system with no additional components to perform active control methods to balance flow across multiple battery packs 120. Thus, the weight of the system may be minimized. Additionally, a passive cooling system may offer a simplified design, reducing manufacturing costs and eliminating points of failure.

[0114] To charge multiple battery packs 120 installed in the wings of the aircraft 100 at a specific charging rate, it is desirable to regulate the battery packs 120 so that their respective temperatures are as close as possible to each other. Therefore, a refrigerant line architecture according to the present disclosure may be designed to maintain a substantially uniform cooling rate by improving refrigerant flow distribution and temperature distribution. Furthermore, it may be desirable to locate the main inlet 1610 and the main outlet 1650 on the same side of the aircraft 100. In other words, the main inlet 1610 and the main outlet 1650 may be located on the same side of the center of the fluid transfer assembly 1601, for example. In yet another way, the main inlet 1610 and the main outlet 1650 may be located closer to a first half of the multiple battery packs 120 than to a second half of the multiple battery packs 120. This may facilitate passenger ingress and egress and enable cooling maintenance operations in parallel with the loading and unloading of passengers and cargo.

[0115] In some embodiments, the main inlet 1610 and the main outlet 1650 may be located adjacent to one another or within a single fixture, such that the supply and return lines of a ground-based refrigerant supply may be attached by a single nozzle attachment. Additionally, in some embodiments, the main inlet 1610 and the main outlet 1650 may be integrally fabricated with other maintenance attachments, such as charging ports, to allow for quick connection of all maintenance utilities. This may allow for easier maintenance operations or allow operations to be performed by autonomous devices. Furthermore, while a ground-based refrigeration supply is mentioned, embodiments of the present disclosure are not limited thereto. For example, in some embodiments, refrigerant and power may be provided by, for example, an air-based refrigerant supply and quick charging station. The air-based refrigerant and charging station may include, for example, a dedicated VTOL or CTOL aircraft configured to pass and / or follow the aircraft 100 in the air to provide refrigerant or charge the aircraft 100. This may, for example, allow for longer continuous flight or provide emergency maintenance services in areas where landing is suboptimal or impossible. Thus, in some embodiments, the main entrance 1610 and main exit 1650 may be provided in a self-docking fixture, along with charging ports and the like, for in-flight attachment and detachment.

[0116] In FIG. 16A , a “U-flow” type cooling system 1600A is provided. The cooling system 1600A may include multiple heat exchangers corresponding to the multiple battery packs 120 and a fluid delivery assembly 1601. In the fluid delivery assembly 1601 of FIG. 16A , the main inlet 1610 and the main outlet 1650 may be located adjacent to each other or on the same side of the aircraft 100. The coolant flow may enter the battery packs 120 in parallel and return to the main outlet 1650, forming a generally “U” shaped structure in the schematic diagram. In some embodiments, the inlet path 1611 and / or the outlet path 1651 may include multiple segments 1612 and 1652 between each branch leading to the heat exchanger inlet 1621 or the heat exchanger outlet 1622, respectively. As discussed further below, the segments may be different sizes to provide balancing flows to each battery pack 120.

[0117] In some embodiments, the U-shaped flow design may comprise a “first-in, last-out” architecture, such that the first battery pack 120 to receive refrigerant from the inlet path 1611 in the inlet flow direction may be the last battery pack 120 to discharge refrigerant to the outlet path 1651 in the outlet flow direction. This results in differences in the path lengths of the individual refrigerant flow loops through each individual battery pack. For example, as seen in FIG. 16A , the refrigerant flowing through the right-most battery pack 120 may follow the shortest loop between the main inlet 1610 and the main outlet 1650, while the refrigerant flowing through the left-most battery pack 120 may follow the longest loop. While this U-shaped flow arrangement allows for the desired placement of the main inlet 1610 and the main outlet 1650 in close proximity to one another, pressure and flow imbalances can occur if not properly considered. Some embodiments of the present disclosure provide systems and methods for passively maintaining a balancing flow while keeping the main inlet 1610 and the main outlet 1650 together.

[0118] FIG. 16B illustrates an alternative “Z-flow” cooling system 1600B. In the Z-flow architecture, the main inlet 1610 may be located on one side of the aircraft 100 and the main outlet 1650 may be located on the other side, forming a generally “Z”-shaped structure in schematic view. The Z-flow refrigerant line architecture 1600B may allow the path length of the refrigerant entering each of the six battery packs 120 to be identical, which may be desirable to optimize flow and temperature distribution. For example, unlike the U-flow design described above, the order of the heat exchanger inlets and outlets in the Z-flow design may comprise a “first-in, first-out” structure, such that along the refrigerant flow path, the first battery pack 120 to receive refrigerant from the inlet path 1611 in the inlet flow direction may be the first battery pack 120 to discharge refrigerant to the outlet path 1651 in the outlet flow direction. This results in substantially the same overall refrigerant flow path length for all battery packs. For example, the battery pack with the shortest distance to the main inlet 1610 may have the longest distance to the main outlet 1650.

[0119] In either the U-flow refrigerant line architecture 1600A or the Z-flow refrigerant line structure 1600B, the flow distribution within the battery packs 120 may be balanced, for example, by varying the inner diameter between different segments of the inlet or outlet paths. For example, the inlet path segment immediately upstream of the right-most (first) battery pack 120 in FIG. 16B may be sized to supply refrigerant to all battery packs 120, while the inlet path segment immediately upstream of the left-most (last) battery pack 120 may be sized to accommodate only one battery pack. Thus, in some embodiments, the inner diameter of the inlet path segments may continuously decrease toward the inlet path. In some embodiments, a reverse arrangement may be configured relative to the outlet path in FIG. 16B . For example, the inner diameter of the outlet path segments may continuously increase toward the outlet path. Thus, a refrigerant flow architecture is provided that uses minimal piping or hose sizes at each segment along the supply and return paths, which may further minimize the wet and dry weight of the cooling system.

[0120] To simultaneously maintain balanced flow while keeping the main inlet 1610 and the main outlet 1650 on the same side of the aircraft 100, FIG. 16C provides a modified “U-flow” cooling system 1600C in which customized flow restrictors 1630 may be positioned along the inlet (and / or outlet) paths between the main inlet 1610 (and / or main outlet 1650) and one or more battery packs 120, respectively. In other words, the coolant may flow through the customized flow restrictors 1630 before entering the corresponding battery pack 120. Each customized flow restrictor 1630 may provide a desired flow resistance to achieve an appropriate amount of coolant flow and balance the flow entering each of the battery packs 120. For example, in the cooling system 1600C, the flow restrictors 1630 may include sized orifices to equalize the pressure drop from the common main inlet 1610 to each battery pack 120. This may ensure uniform flow distribution of the coolant to each battery pack 120. In this way, the cooling architecture may function efficiently without the need to provide multiple different piping diameters or equivalent path lengths.

[0121] 16D illustrates a modified "Z-flow" cooling system 1600D that adds an in-flight return path 1640. The in-flight return path 1640 may include an additional leg of an outlet path 1651 located within the wing of the aircraft 100, for example. By providing the in-flight return path 1640, the main inlet 1610 and the main outlet 1650 may be located on the same side of the aircraft 100, while still balancing the flow distribution within the battery pack 120, for example, by using a Z-flow architecture. Similar to the Z-flow cooling system 1600B, in some embodiments, this design may include different sized piping segments or multiple different sized orifices.

[0122] Additionally, by providing segments with continuously increasing / decreasing inner diameters, using flow restrictors, and / or providing equal path lengths for each loop as described above, embodiments of the present disclosure may achieve a symmetrical architecture configured to provide uniform flow characteristics regardless of whether the coolant is circulating in the forward or reverse direction. For example, in some embodiments, the coolant architecture may be configured for symmetrical flow in both directions. This may allow for improved cooling uniformity by alternately circulating the coolant in different directions, as discussed further below, for example, with respect to FIG. 20.

[0123] 16A-16D and other figures may be referred to by terms such as "inlet" and "outlet," it should be understood that in some embodiments, such elements may be configured to allow refrigerant to flow in either direction. By way of example, in some embodiments, the main inlet 1610 may function as a first main inlet, the main outlet 1650 may function as a second main inlet, the heat exchanger inlet 1621 may function as a first heat exchanger inlet, and the heat exchanger outlet 1622 may function as a second heat exchanger inlet. Furthermore, in some embodiments, the entire cooling system between the first and second main inlets is free of check valves or other one-way flow fixtures to allow such reverse flow operation. Additionally, to ensure symmetrical flow characteristics in the forward and reverse flow directions, in some embodiments, the inlet path 2011 and the outlet path 1651 may have substantially equal lengths.

[0124] FIG. 17 illustrates an exemplary battery cooling system 1700 based on the cooling system of FIG. 16C , consistent with some embodiments of the present disclosure. As previously described, the aircraft 100 may include a common main inlet 1710 and a common main outlet 1750 for receiving and discharging coolant. The coolant may flow through respective flow restrictors 1730 added to tubing associated with the battery packs before flowing in parallel to each of the battery packs 120. In some embodiments, each flow restrictor 1730 may have a different length and / or orifice size depending on desired flow characteristics. For example, the orifice size may increase along the inlet flow direction to compensate for pressure drops that occur as the coolant branches toward the upstream battery packs 120. As an illustrative, non-limiting example, the flow restrictors 1730a-f may have orifice sizes in the inlet flow direction of, for example, 3.7 mm, 4.1 mm, 4.65 mm, 6.8 mm, 10 mm, and 12.5 mm, respectively, and tubing diameters of, for example, approximately 20 mm or 0.5 inches. The flow restrictor may be located upstream or downstream of the battery pack 120. For example, embodiments of the flow restrictor may be compatible with the reverse flow operation disclosed herein.

[0125] In some embodiments, the total flow rate of the cooling system may be designed to be, for example, approximately 60 liters per minute (LPM). That is, for each of the six battery packs 120, the individual flow rate may be approximately 10 LPM. However, the present disclosure is not limited in this regard. In general, any suitable total flow rate may be divided substantially equally among the battery packs 120. In some embodiments, for example, the flow rate through the heat exchanger of each battery pack 120 may deviate from the average value of all the battery packs 120 by, for example, 5%, 2%, 1%, 0.5% or less. The cooling system may also be robust to errors in refrigerant flow and distribution. For example, continuing the example above, if uneven distribution within the cooling system causes refrigerant to be distributed among the six battery packs 120 at twice the flow rate of 120 LPM, each pack may receive an average flow rate of 20 LPM. For example, the differential pressure between the battery packs 120 may be approximately 20 psi when the flow rate is 10 LPM and approximately 40 psi when the flow rate is 20 LPM.

[0126] In some embodiments, instead of or in addition to providing flow restrictors 1630a-f, the various segments 1712 and 1752 of the inlet and outlet paths of cooling system 1700 may have different inner diameters, as described above. For example, the inlet path may include six segments 1712a-f, and the outlet path may include six segments 1752a-f. The inlet segments 1712a-f may be configured to provide a continuously decreasing flow rate in the inlet flow direction. Alternatively or additionally, the outlet segments 1752a-f may be configured to provide a continuously increasing flow rate in the outlet flow direction.

[0127] Reference is made to FIG. 18 , which illustrates an exemplary home system 1800 consistent with certain embodiments of the present disclosure. It is understood that batteries and / or battery systems used in demanding applications on aircraft, such as high-voltage batteries manufactured primarily for powering eVTOL aircraft, will experience progressive degradation of battery characteristics, rendering the battery and / or battery system unusable for its primary application over time. Typical parameters indicative of this degradation include a decrease in energy storage capacity below a threshold, an increase in temperature rise under high-stress (current, power) conditions encountered during aircraft use, an increase in the internal resistance / impedance of the battery or battery system, and a reduction in power delivery capacity to levels below the requirements of the aircraft's host devices. However, such high-voltage batteries or battery systems may continue to be useful in other secondary applications, including, but not limited to, home power systems, rural energy storage systems, and other backup power systems, including bulk power grid and home grid backup systems. For example, secondary uses of such batteries and / or battery systems may include powering a home, where the reused battery or battery system may be used as an energy source to power the home, and the battery or battery system may be directly connected to a home system inverter and power source.

[0128] In some embodiments of the present disclosure, a high-voltage battery originally used in an aircraft can be reliably reused for a secondary use. In some embodiments, the high-voltage battery can be a battery for an eVTOL aircraft. In some embodiments, the high-voltage battery may be originally designed from the beginning for its original use and for reuse in the secondary use. For example, the high-voltage battery may be originally designed so that minimal modification may be required before reuse in the secondary use. As one example, the high-voltage battery may be originally designed with a removable high-voltage junction box including a battery management system, such that modification to the secondary use may include replacing the original junction box with another junction box (including a different battery management system) configured for the secondary use. As another example, the high-voltage battery may be designed to operate in conjunction with an inverter compatible with both the original use and the secondary use. For example, the inverter's voltage, current, power, temperature, humidity, and other ratings may be configured so that the inverter is compatible with both the original use and the secondary use. As another example, a high-voltage battery may be configured with a battery interface that is compatible with the interface of both the original application (e.g., eVTOL aircraft) and the secondary application (e.g., home energy storage system) such that the battery interface does not need to be changed when the high-voltage battery is reused for the secondary application.

[0129] In some embodiments, the secondary use may be installed in a home system (i.e., a residential power system) and supply power to the home system. In some embodiments, the home system may be a utility system that powers various subsystems of a home, residential complex, commercial facility, or industrial facility. In some embodiments, the home system may include an electronic switching system (ESS). As shown in the example of FIG. 18 , home system 1800 may include a recycled battery 1810, where recycled battery 1810 is coupled to a battery interface 1815, where battery interface 1815 is coupled to an inverter 1820, where inverter 1820 is coupled to one or more line filters 1825. In some embodiments, battery 1810, battery interface 1815, inverter 1820, and one or more line filters 1825 may comprise an electronic switching system 1860 of home system 1800. In some embodiments, battery interface 1815 may be a DC / DC interface. In some embodiments, such an exemplary electronic switching system 1860 may be connected to one or more home electrical lines 1840 via a switching network (i.e., switchgear) 1830. In some embodiments, the switching network 1830 may be configured to allow power to be fed to the home electrical lines 1840 either via the electronic switching system 1860 or via an external power grid 1850.

[0130] Reference is now made to FIG. 19. FIG. 19 illustrates an exemplary electronic switching system 1900 for the home system 1800 shown in FIG. 18. As shown in the example of FIG. 19, the electronic switching system 1900 may include a recycled HV battery and battery interface 1920 (the battery interface including a battery management system (BMS)), and power electronics equipment 1910 including an inverter and one or more line filters. In some embodiments, such an exemplary electronic switching system 1900 may be connected to one or more home electric lines via a switching network (i.e., switchgear). In some embodiments, the switching network may be configured to allow power to be fed into the home electric line either through the electronic switching system or an external power grid. In some embodiments, the inverter within the power electronics equipment 1910 of the electronic switching system 1900 may be configured to have minimal total harmonic distortion (THD), thereby requiring minimal filtering to connect the electronic switching system to a utility bus and thereby providing minimal harmonic distortion in any power returned to the electronic switching system 1900 and the external power grid connected to the home system.

[0131] Continuing to refer to FIGS. 18 and 19 , in some embodiments, the repurposed batteries may be high-power (i.e., HV) batteries with high-power cells that may enable increased power applications in the home electrical system 1840. As an example, one or more of such HV batteries with high-power cells may be utilized for DC fast charging (DCFC) applications in the home. In some embodiments, the DCFC applications may include DCFC of electric vehicles (EVs) at rates significantly greater than currently possible. For example, two high-power batteries with high-power cells described herein may be used in a home system to charge an EV in approximately 40-80% of the time typically required by a V3 Supercharger and at approximately 25 times the rate of existing home EV charging stations. As another example, to further support DCFC applications via the electronic switching system 1900, custom DC / DC battery interfaces may be required to support DCFC capabilities. As a result, in some embodiments, the custom DC / DC battery interface may be minimized in size, optimized for efficiency, and / or include control logic, sensors, and / or communication interfaces that further support the vehicle's DCFC, for example, via the home electronic switching system 1900.

[0132] 19 , in further embodiments where the recycled battery is a high-power battery with high-power cells, the recycled battery may further have a low resistance. In some embodiments, the recycled battery may further generate a lower amount of heat compared to batteries in existing home electronic switching systems. As a result, in some embodiments, the recycled battery may be utilized as a heat sink for the inverter. As a further result, in some embodiments, such an inverter may not require any additional forced cooling for operation. As a still further result, in some embodiments, an electronic switching system including a high-power battery with high-power cells may be a quieter and / or more compact system.

[0133] 18 and 19 , in some embodiments, the home system includes parallel batteries 1810, at least one of which is a recycled battery. In some embodiments, adding additional capacity to the electronic switching system 1860 by including additional batteries connected in parallel to the first battery 1810 of the electronic switching system 1860 may not require additional equipment. For example, in some embodiments, the electronic switching system 1860 may include a master battery that further includes an interface, an inverter, and one or more line filters, and the additional batteries may be connected in parallel to the master battery. In some embodiments, including batteries connected in parallel to the master battery may meet a specific capacity need for a particular home.

[0134] In some embodiments, a recycled battery may include most of the sensors necessary to function properly within electronic switching system 1900 and / or home system 1800. For example, a recycled battery for use in electronic switching system 1900 and / or home system 1800 may only require replacing the aircraft-specific high voltage junction box (HVJB) with a home system-specific HVJB. In some embodiments, a recycled battery for use in electronic switching system 1900 and / or home system 1800 may only require a positive battery connector, a negative battery connector, and a 5x cell monitoring unit (CMU) connector.

[0135] In some embodiments, a common HVJB may be installed in the battery from the initial build of the battery for its primary use. In such embodiments, the common HVJB may eliminate the requirement to replace the HVJB with a home system-specific HVJB before the secondary use of the battery. Because excess processing and / or storage capacity is needed for aircraft batteries, such batteries may also have capacity that can be preloaded to configure home system compatibility and / or functionality during their initial build.

[0136] FIG. 20 is a schematic diagram of a cooling system 2000 for battery cooling in an HV cooling distribution subsystem consistent with some embodiments of the present disclosure. As previously described, the aircraft 100 may include, for example, six battery packs 120. Each battery pack 120 may include an inlet and an outlet for circulating a refrigerant through the heat exchanger assemblies of the battery pack 120 to provide battery cooling. Similar to the embodiments of FIGS. 16-17 described above, FIG. 20 proposes a refrigerant line architecture with refrigerant inlets and outlets for circulating refrigerant in parallel to the six battery packs 120 installed in the wings of the aircraft 100, minimizing both the dry and wet weights of the cooling system.

[0137] 20 embodiment, the cooling system 2000 may include a heat exchanger assembly corresponding to each of the battery packs 120 and a fluid delivery assembly 2001 coupled to the heat exchanger assembly. Each heat exchanger assembly may include a first heat exchanger inlet / outlet 2021 and a second heat exchanger inlet / outlet 2022 configured to receive a heat transfer fluid (i.e., refrigerant) from or discharge a heat transfer fluid to the fluid delivery assembly 2001. The fluid delivery assembly 2001 may be configured to circulate the heat transfer fluid in parallel with the heat exchanger assemblies of the battery packs 120 in a dual U-flow manner.

[0138] For example, the dual U-flow system may include a first main entrance 2010 connected to a first trunk 2013 and a second main entrance 2050 connected to a second trunk 2053. The first trunk and the second trunk may each branch, for example, at a central location of the aircraft 100, into a first U-flow loop that serves a first set of battery packs 120 in a first wing of the aircraft 100 and a second U-flow loop that serves a second set of battery packs 120 in a second wing of the aircraft 100. The first trunk 2013 and the second trunk 2053 may each branch. For example, the first trunk 2013 may branch into a first branch 2014 and a third branch 2015. The second trunk 2053 may branch into a second branch 2054 and a fourth branch 2055. In some embodiments, to ensure symmetrical characteristics in the forward and reverse flow directions, the first main line 2013 and the second main line 2053 may have substantially equal lengths, the first branch 2014 and the second branch 2054 may have substantially equal lengths, or the third branch 2015 and the fourth branch 2055 may have substantially equal lengths.

[0139] As indicated by the arrows in Figure 20, the first trunk 2013 operates as an inlet trunk, and the first branch 2014 and the third branch 2015 operate as inlet paths for their respective U-flow loops. Similarly, the second trunk 2053 operates as an outlet trunk, and the second branch 2054 and the fourth branch 2055 operate as outlet paths for their respective U-flow loops. However, in some embodiments, as described above with respect to Figures 16A-17, the cooling system 2000 may be configured as a symmetrical system capable of circulating refrigerant flow in either direction, such that either set of trunks and branches may be used as inlet or outlet lines.

[0140] Providing refrigerant flow in alternating directions during continuous circulation may improve cooling efficiency or uniformity. For example, in a given cooling operation, the flow direction may be reversed one or more times. For example, the flow direction may be changed midway through a cooling cycle or at other regular or irregular intervals. In some embodiments, the refrigerant may flow primarily in one direction, periodically interrupted by back-and-forth flow. The external refrigerant supply may be configured to control such changes in flow direction, for example, by operation of a pump, flow exchange valve, etc.

[0141] The first and second main inlets of the transport assembly 2001 may be integrally located at the same location on the aircraft 100. For example, the first and second main inlets may be located at the top, bottom, or side of the fuselage of the aircraft 100, or within a wing. For example, a location at the top of the fuselage, e.g., between or behind the wings, may advantageously reduce the overall length of the hoses due to proximity to the battery packs 120. However, a location at the side or bottom of the fuselage may allow for easier connection access. In some embodiments, when the first and second main inlets are located in an asymmetrical location, such as on one side of the fuselage, the inlet and outlet trunks may nevertheless be configured to maintain substantially equal lengths before splitting into the first and second U-shaped flow loops. For example, the first and second trunks may terminate in their respective branches at a central location on the aircraft 100. In this manner, symmetrical flow path lengths may be maintained while providing asymmetrical connections to external coolant sources. As described above, the first main inlet and second main inlet may be configured as a single connection port so that an external coolant supply source may provide a flow of balancing coolant to all battery packs of the aircraft 100 simultaneously through a single connection.

[0142] In some embodiments, as described above, a customized flow restrictor 2030 may be disposed between the branch and one of the first heat exchanger inlet / outlet 2021 or the second heat exchanger inlet / outlet 2022, respectively. The flow restrictor 2030 may be configured to balance the heat transfer fluid entering the heat exchanger assembly of the battery pack 120, as described above.

[0143] In some embodiments, as an example, the flow rate of the coolant system may be designed to be approximately 30 liters per minute (LPM) within each U-flow loop. Thus, in a configuration where each U-flow loop serves three battery packs 120, the flow rate may be approximately 10 LPM for each battery pack 120. However, the present disclosure is not limited in this respect. In general, the flow rate through each battery pack 120 may deviate from the average value across all battery packs 120, as discussed above, by, for example, 5%, 2%, 1%, 0.5%, or less.

[0144] 21A and 21B illustrate an example design of a cooling system 2100 consistent with some embodiments of the present disclosure. The cooling system 2100 may correspond, for example, to the cooling system 2000 of FIG. 20. As shown in FIG. 21A, a fluid delivery assembly 2101 may provide battery coolant to each battery pack 120. The fluid delivery assembly may include a first main inlet / outlet 2110 and a first trunk 2111, and a second main inlet / outlet 2150 and a second trunk 2151. The first trunk 2113 and the second trunk 2153 may each split into a first branch, a second branch, a third branch, and a fourth branch, as described above. 21A, the first and second main ports of the fluid delivery assembly 2101 may be located, for example, at the top of the fuselage at mid-wing span of the aircraft 100 to achieve a more balanced cooling distribution with three battery packs 120 located on one side and three battery packs 120 located on the other side. Alternatively or additionally, as described above, the first and second main ports may be located, for example, on one side of the fuselage or on one wing.

[0145] FIG. 21B shows a further view of the cooling system 2100 according to an embodiment of the present disclosure. The view of FIG. 21B may be considered an inverted view of FIG. 21A, showing the underside of the cooling system 2101. FIG. 21B shows details of various conduit sections of the fluid delivery system 2101 and their connection to the heat exchanger 2160. Here, for clarity, the heat exchanger is omitted from the first U-flow loop at the bottom of the figure. The flow path segment 2112 of the first branch 2114 from the first trunk 2113 may optionally be connected to, for example, the first heat exchanger inlet / outlet 2121 via a customized flow restrictor (not shown). Similarly, the flow path segment 2152 of the second branch 2154 from the second trunk 2153 may be connected to, for example, the second heat exchanger inlet / outlet 2122. In some embodiments, the conduit portions of the fluid delivery assembly 2101 may be constructed of various materials. For example, as shown in FIG. 21B , the first trunk 2113 and the second trunk 2152 may comprise a rigid tubing material, such as aluminum or other high-strength, lightweight material. Additionally, the first segment 2112a of the first branch 2114 may be constructed of the same or a similar material. The subsequent segment 2112 may be constructed of a different material, such as a flexible hose. Similarly, the second segment 2152a of the second branch 2154 may be constructed of the same or a similar material. The subsequent segment 2152 may be constructed of a different material, such as a flexible hose. Constructing downstream segments of branches in the fluid delivery assembly may allow for easy installation, maintenance, and replacement. In some embodiments, other materials may be used. For example, instead of rigid tubing, the channels may be integrally formed into the body of the aircraft 100. Instead of hoses, for example, polyethylene or other semi-rigid tubing may be provided. In some embodiments, all conduits in the fluid delivery assembly may be constructed of the same material.

[0146] FIG. 22A illustrates an exemplary heat exchanger assembly 2200 for battery cooling in an HV cooling distribution subsystem, consistent with certain embodiments of the present disclosure. FIG. 22B illustrates a top view of the heat exchanger assembly 2200 illustrated in FIG. 22A , consistent with certain embodiments of the present disclosure. The design of the heat exchanger assembly 2200 illustrated in FIGS. 22A and 22B enables bottom-up cooling of multiple battery cells within the battery pack 120. The heat exchanger assembly 2200 may have a structure and dimensions configured to minimize the dry and wet weight of the battery pack 120. In the embodiment of FIGS. 22A and 22B , the heat exchanger assembly 2200 is integrated into a thermal management system to ensure maximum cooling capacity without exceeding a desired system-wide pressure limit (e.g., less than 100 psi), eliminating the need for special or customized connectors. Furthermore, by minimizing thermal gradients between cells, this design may also extend the battery life of the battery pack 120.

[0147] In particular, in the heat exchanger assembly 2200, a plurality of dimples 2210 are incorporated on the heat exchanger plate (i.e., cooling plate) 2220 to function as turbulence generators to improve the heat transfer efficiency of the heat exchanger assembly 2200. For example, the dimples 2210 may comprise formed depressions within the cooling channels 2222 of the heat exchanger assembly 2200 configured to create a calculated distortion within the interior volume to disrupt laminar coolant flow, resulting in a more uniform distribution of surface contact between the coolant molecules and the inner walls of the cooling channels. In some embodiments, the heat exchanger plate 2220 may be hydroformed and laser welded. The dimples 2210 may be created during the hydroforming process or stamped thereafter, for example. The heat exchanger plate 2220 may be, for example, an aluminum plate or another lightweight, strong, thermally conductive material suitable for cost-effective manufacturing. This approach may enable feasible mass production and effective quality control.

[0148] Specifically, the heat exchanger plate 2220 may include a plurality of cooling channels 2222, a first manifold channel 2224, and a second manifold channel 2226. The first manifold channel 2224 may be coupled to a first heat exchanger port 2240 for receiving or discharging a heat transfer fluid (i.e., a refrigerant), and the second manifold channel 2226 may be coupled to a second heat exchanger port 2250 for receiving or discharging the heat transfer fluid. The plurality of cooling channels 2222 may be coupled between the first manifold channel 2224 and the second manifold channel 2226 and arranged such that the heat transfer fluid flows through the cooling channels 2222.

[0149] Each cooling channel 2222 may include a plurality of dimples 2210 configured to provide turbulent flow of heat transfer fluid through the cooling channel 2222. The arrangement of the dimples 2210 may be configured to meet the particular needs of, for example, an electric aircraft. For example, it may be desirable for the dimples 2210 to be symmetrical in both shape and arrangement along the flow direction so that their heat transfer characteristics are the same in either flow direction. This may improve performance of the bidirectional cooling operation described above.

[0150] Additionally, dimples 2210 may be configured to introduce localized turbulence into the coolant flow without excessively disrupting laminar flow. For example, if dimples 2210 cause excessive separation of the coolant flow on the dorsal side of dimples 2210, overall cooling efficiency and uniformity may be reduced. Purging the cooling channels may also be more difficult, which may be desirable in some embodiments to minimize the weight of the electric aircraft. However, creating complete turbulence within the cooling channels is undesirable because the resulting pressure drop would make it difficult, impractical, or impossible to achieve high coolant flow rates (e.g., 100 or 200 liters per minute) that may be desirable in some embodiments. Finally, it is desirable to create a dimple geometry that is cost-efficient to manufacture while still achieving these benefits.

[0151] Thus, in some embodiments, as shown with respect to FIGS. 24A-24B, the dimples 2210 may be staggered to improve heat transfer and achieve optimal flow characteristics. This can provide significant advantages over dimples located in the center of the cooling channel. For example, centrally located dimples located periodically along the length of the channel may cause excessive flow separation at low pitches (as seen in dimple arrangement 2401 on the left side of FIG. 24A) compared to high pitches (as seen in dimple arrangement 2402 on the right side of FIG. 24A). By staggering dimples on opposite sides of the channel (as seen in dimple arrangement 2403 on the right side of FIG. 24B), the coolant may follow a tortuous path that does not cause excessive flow separation, even with a dimple configuration with a similar pitch to dimple arrangement 2401. For example, the dimples may be configured to generate turbulent coolant flow, with each dimple introducing localized turbulence into an otherwise laminar flow downstream of the dimple without completely changing the flow from laminar to turbulent. In some embodiments, the channels may be configured to generate flow with a Reynolds number (based on the channel hydraulic diameter without dimples), for example, between 100 and 2000.

[0152] Additionally, the dimples 2210 may be advantageously configured with flow-symmetric shapes. For example, a staggered semicircular shape may provide desired flow characteristics. Furthermore, the dimples 2210 may have substantially flat, forward-sloping faces when viewed from either direction of flow within the cooling channels 2222. For example, in some embodiments, the dimples 2210 may comprise a triangular or trapezoidal shape. The dimples 2210 may be recessed from the top side of the heat exchanger plate 2220 to the top surface of the cooling channels 2222 when viewed from their installation direction. This may help minimize deposition of refrigerant droplets during purging and enable cost-effective manufacturing. Furthermore, the dimples 2210 within each cooling channel 2222 may be staggered or arranged side-to-side. For example, in some embodiments, as further illustrated with respect to Figures 23A-23B below, any two adjacent dimples 2210 may be recessed from opposite sides of the top (or bottom) surface of the cooling channel 2222 to achieve serpentine turbulence without causing excessive flow separation as described above. Alternatively or additionally, the dimples 2210 may be recessed from the side of the cooling channel 2222, for example.

[0153] In some embodiments, the cooling channels 2222 disposed within the heat exchanger assembly 2200 may have a relatively small channel height compared to their channel width when viewed from an installation direction. For example, if the cooling channel height is, for example, approximately 1.5 mm, the channel width may be approximately 18.2 mm. In such a case, the manifold sections (e.g., manifold channels 2224 and 2226) may have a channel height of approximately 5 mm and a channel width of approximately 25 mm. In other words, the height of the first manifold channel 2224 or the second manifold channel 2226 may be greater than the height of the cooling channel 2222. The channel height may be designed to minimize both the dry and wet weight of the heat exchanger assembly 2200 and to allow for selective purging and replenishing of the refrigerant as needed during ground maintenance of the aircraft 100.

[0154] In the heat exchanger assembly 2200, triangular dimples 2210 are added to increase heat transfer efficiency and minimize temperature gradients within the battery pack 120. As will be appreciated by one skilled in the art, in various embodiments, the triangular dimples 2210 may be formed using various manufacturing methods.

[0155] In this embodiment, the heat exchanger assembly 2200 may include symmetrical inlets and outlets 2240 and 2250 for receiving and returning refrigerant, respectively. That is, the heat exchanger inlets and outlets 2240 and 2250 may be configured to either receive a heat transfer fluid or discharge a heat transfer fluid. Specifically, the symmetrical design for the heat exchanger inlets and outlets 2240 and 2250 of the heat exchanger assembly 2200 may enable the implementation of a reverse flow strategy, as described above. For example, the direction of refrigerant flow may be reversed as needed, for example, by switching the inlets and outlets, operating an external pump, or by using a flow exchange valve. For example, the heat exchanger plate 2220 may be configured to receive heat transfer fluid from the first heat exchanger inlet / outlet 2240 and discharge the heat transfer fluid to the second heat exchanger inlet / outlet 2250 during a first operating period, and to receive heat transfer fluid from the second heat exchanger inlet / outlet 2250 and discharge the heat transfer fluid to the first heat exchanger inlet / outlet 2240 during a second operating period before or after the first operating period.

[0156] 22C illustrates an exemplary refrigerant flow path within a heat exchanger assembly 2200 for battery cooling in an HV cooling distribution subsystem, consistent with certain embodiments of the present disclosure. As shown in FIG. 22C, in accordance with the embodiments of FIGS. 22A and 22B above, for example, heat exchanger assembly 2200 may employ a "10P3S" or similar topology. Such a topology may include ten cooling units (e.g., cooling unit 2230) connected in parallel, with each cooling unit comprising three longitudinal cooling channel paths (e.g., cooling channels 2222a, 2222b, and 2222c) connected in series to form an "S-curve" shape.

[0157] In other words, the heat exchanger plate 2220 may be configured to organize the cooling channels (e.g., cooling channels 2222a, 2222b, and 2222c) into cooling units 2230. The cooling units 2230 may be coupled in parallel between the first manifold channel 2224 and the second manifold channel 2226 such that the cooling fluid is substantially evenly distributed among the ten separate cooling unit paths, with each subdivision of the cooling fluid passing through all three cooling channels 2222a, 2222b, and 2222c of their respective cooling units 2230. In some embodiments, other related topologies, such as 6P6S, 7P5S, etc., may be applied. In general, the heat exchanger assembly 2220 may comprise multiple cooling units arranged in parallel between the first manifold channel 2224 and the second manifold channel 2226, each cooling unit including an odd number of alternating cooling channel directions in series to achieve inlet and outlet at opposing manifold channels.

[0158] In some embodiments, the total volume of refrigerant in the heat exchanger assembly 2200 (i.e., not including the volume of the conduits in the fluid delivery assemblies) can be approximately 0.7 L. The 10P3S topology can be configured to balance the flow distribution and pressure drop between the parallel cooling channels. Furthermore, the three cooling channels connected in series in the 3S crossflow design shown in FIGS. 22A-22C can reduce thermal gradients that occur as the temperature of the refrigerant increases. As shown in FIG. 22C, in forward flow operation, after the heat exchanger assembly 2200 receives refrigerant from an inlet (e.g., heat exchanger inlet / outlet 2240), the refrigerant flows into a first manifold section (e.g., first manifold channel 2224) and is then substantially evenly distributed to multiple cooling units 2230. Each cooling unit 2230 can include, for example, three cooling channels connected in series (e.g., cooling channels 2222a, 2222b, and 2222c). The cooling channels in the heat exchanger assembly 2200 may be aligned parallel to one another. In some embodiments, the center-to-center distance between two adjacent cooling channels may be, for example, about 4 mm, 6 mm, 8 mm, 10 mm, 15 mm, 20 mm, or more.

[0159] As shown, within each cooling unit 2230, the flow direction may change upon entering the next cooling channel. That is, the heat transfer fluid may flow in opposite directions within any two consecutive cooling channels within the same cooling unit 2230. For example, in forward flow operation, the refrigerant within cooling channels 2222a and 2222c flows to the right, as seen in FIG. 22C, while the refrigerant within cooling channel 2222b flows to the left. After exiting the last cooling channel 2222c within cooling unit 2230, the refrigerant from the ten cooling units flows into a second manifold section (e.g., second manifold channel 2226) and exits through an outlet (e.g., heat exchanger inlet / outlet 2250, as seen in FIG. 22A).

[0160] On the other hand, in reverse flow operation, the heat exchanger assembly 2200 receives refrigerant through an inlet (e.g., heat exchanger port 2250). The refrigerant then flows into a second manifold section (e.g., second manifold channel 2226) and is distributed in parallel to the cooling units 2230. The refrigerant in cooling channels 2222a and 2222c in each cooling unit 2230 flows to the left, while the refrigerant in cooling channel 2222b in each cooling unit 2230 flows to the right. After exiting the last cooling channel 2222a in cooling unit 2230, the refrigerant from all ten cooling units flows into a first manifold section (e.g., first manifold channel 2224) and exits through an outlet (e.g., heat exchanger port 2240 as seen in FIG. 22A ).

[0161] FIG. 23A illustrates an enlarged portion of a cooling channel 2300 in the heat exchanger assembly 2200 of FIG. 22A , consistent with certain embodiments of the present disclosure. FIG. 23B illustrates a cross-sectional view taken along line AA of FIG. 23A , consistent with certain embodiments of the present disclosure. As shown in FIGS. 23A and 23B , a plurality of triangular dimples 2310 are arranged in a staggered pattern, with two adjacent dimples 2310 recessed from the top surface of the cooling channel 2300 and positioned on opposite lateral sides. In various embodiments, the dimensions of the cooling channel 2300 and dimples 2310 can be designed to accommodate specific implementations of embodiments of the present disclosure based on actual needs to optimize overall heat transfer efficiency.

[0162] For example, design parameters that affect the coolant flow characteristics, and therefore the heat transfer characteristics, in the cooling channel 2300 may include dimensional parameters such as the width 2320 of the cooling channel 2300, the height 2330 of the cooling channel 2300, the distance 2340 between two adjacent dimples 2310 (e.g., the distance between the apexes of two adjacent dimples 2310), the width 2312 of the dimple 2310, the height 2314 of the dimple 2310, the base angle of the dimple 2316 between the sidewall of the cooling channel and the substantially flat, forward-sloping face of the dimple, the length 2318 of the apex region of the dimple 2310, etc. The optimal values ​​of such parameters may depend on the properties of the applied coolant, such as viscosity, heat transfer coefficient, pressure, flow rate, etc. In some embodiments, the dimensions of the dimple 2310 (e.g., width 2312, height 2314, length 2318, etc.) may also be expressed using dimensionless ratios based on the dimensions of the cooling channel 2300, including the width 2320 and height 2330.

[0163] The cooling architectures disclosed in various embodiments of the present disclosure may offer several beneficial advantages and are suitable for cooling distribution systems on eVTOL aircraft. For example, the proposed design reduces the mass of the heat exchanger assembly and the required fluid mass in the cooling system. The improved flatness of the design also correspondingly simplifies the manufacturing process, improves overall thermal performance, and can meet critical cell thermal requirements, such as low temperature gradients across each pack and maximum temperatures desired for safety and performance reasons.

[0164] Embodiments of the present disclosure may be further described by the following clauses. Article 1. A cooling system for an electric aircraft, comprising: a plurality of heat exchanger assemblies corresponding to a plurality of battery packs included in the electric aircraft, each heat exchanger assembly including a first heat exchanger inlet and a second heat exchanger inlet configured to receive a heat transfer fluid or discharge the heat transfer fluid; a fluid transport assembly coupled to the heat exchanger assembly, the fluid transport assembly including a first main inlet and a second main inlet, and configured to circulate the heat transfer fluid in parallel with the heat exchanger assembly, the first main inlet and the second main inlet being disposed at the same position on the electric aircraft, the fluid transport assembly comprising: The cooling system further comprises a plurality of conduit segments configured to balance the heat transfer fluid entering each of the heat exchanger assemblies. Clause 2. The cooling system of clause 1, wherein the plurality of conduit segments comprises a plurality of flow restrictors configured to balance the heat transfer fluid flowing into each of the heat exchanger assemblies. Clause 3. The cooling system of clause 2, wherein each of the plurality of flow restrictors comprises a sizing orifice configured to ensure substantially equal inlet pressures at each heat exchanger assembly to achieve uniform flow distribution of the heat transfer fluid. Clause 4. A cooling system as described in any one of clauses 1 to 3, wherein each of the plurality of conduit segments includes a different inner diameter configured to balance the heat transfer fluid flowing into each of the plurality of heat exchanger assemblies. Clause 5. The cooling system of any one of clauses 1 to 4, wherein each heat exchanger assembly is: 1. A heat exchanger plate comprising: a first manifold channel coupled to the first heat exchanger port; a second manifold channel coupled to the second heat exchanger inlet / outlet. Clause 6. A cooling system according to clause 5, comprising: a plurality of cooling channels coupled between the first manifold channel and the second manifold channel, the plurality of cooling channels arranged to allow the heat transfer fluid to flow through the cooling channels, each cooling channel comprising a plurality of dimples configured to provide turbulent flow of the heat transfer fluid through the cooling channel, each of the plurality of dimples comprising a symmetrical contour. Clause 7. The cooling system of clause 5 or 6, wherein the symmetrical contour comprises a first substantially flat, forward-inclined surface when viewed from a first flow direction of the heat transfer fluid and a second substantially flat, forward-inclined surface when viewed from a second flow direction of the heat transfer fluid opposite to the first flow direction. Clause 8. The cooling system of any one of clauses 5 to 7, wherein the symmetrical contour comprises one of a triangle and a trapezoid. Clause 9. A cooling system as described in any one of clauses 5 to 8, wherein the heat exchanger plate is configured to organize the plurality of cooling channels into a plurality of cooling units, and the plurality of cooling units are coupled in parallel between the first manifold channel and the second manifold channel. Clause 10. A cooling system as described in any one of clauses 5 to 9, wherein the heat exchanger plate is configured to receive the heat transfer fluid from the first heat exchanger inlet / outlet and discharge the heat transfer fluid to the second heat exchanger inlet / outlet during a first operating period, and to receive the heat transfer fluid from the second heat exchanger inlet / outlet and discharge the heat transfer fluid to the first heat exchanger inlet / outlet during a second operating period before or after the first operating period. Clause 11. The cooling system of any one of clauses 5 to 10, wherein the heat exchanger plates comprise aluminum. Clause 12. The cooling system of any one of clauses 5 to 11, wherein the plurality of dimples in each cooling channel are arranged in a staggered pattern such that adjacent dimples in the flow direction extend from opposite sides of the cooling channel. Clause 13. The cooling system of clause 12, wherein the adjacent dimples in the flow direction are recessed from laterally opposite sides of a top or bottom surface of the cooling channel. Clause 14. The cooling system of clause 12, wherein the adjacent dimples in the flow direction are recessed from laterally opposite sides of the cooling channel. Clause 15. The cooling system of any one of clauses 1-14, wherein the heat transfer fluid comprises ethylene glycol water (EGW), ethylene glycol, polyethylene glycol, water, or any combination thereof. Clause 16. The cooling system of any one of clauses 1 to 15, wherein the plurality of conduit segments of the fluid transport assembly are: a first tube and a plurality of first hoses coupled to the first tube; The cooling system further comprises a second tube and a plurality of second hoses coupled to the second tube. Clause 17. A cooling system as described in clause 16, wherein the first tube has a first main line leading from the first main inlet / outlet, the plurality of first hoses have first branches connected to the first main line, the second tube has a second main line leading from the second main inlet / outlet, and the plurality of second hoses have second branches connected to the second main line. Clause 18. The cooling system of clause 17, wherein a plurality of flow restrictors are coupled to the first hose or the second hose. Clause 19. The cooling system of any one of clauses 16 to 18, wherein the first tube and the second tube comprise aluminum. Clause 20. The cooling system according to any one of clauses 1 to 19, wherein the first main inlet and the second main inlet are located on a top side of the electric aircraft. Clause 21. The cooling system of any one of clauses 1 to 20, wherein the fluid transport assembly comprises: a first trunk path leading from the first main entrance to a first branch and a third branch; A second trunk line leading from the second main entrance to a second branch and a fourth branch, A cooling system wherein the first trunk and the second trunk are substantially equal in length. Clause 22. The cooling system of clause 21, wherein the first main inlet and the second main inlet are located closer to the first branch than to the second branch. Article 23. A power supply system for an aircraft, comprising: a plurality of battery packs configured to be installed within a wing of the aircraft, each battery pack configured to be connected to an electric engine of the aircraft to provide electrical power; a plurality of heat exchanger assemblies respectively connected to battery pack enclosures of the plurality of battery packs; a fluid delivery assembly coupled to the plurality of heat exchanger assemblies and configured to circulate a heat transfer fluid through the heat exchanger assemblies; each heat exchanger assembly comprising a first heat exchanger port and a second heat exchanger port configured to receive or discharge the heat transfer fluid, and a heat exchanger plate comprising a plurality of dimples configured to provide turbulent flow of the heat transfer fluid; each of the plurality of dimples includes a symmetrical contour having a first substantially flat, forward-sloping surface when viewed in a first flow direction of the heat transfer fluid and a second substantially flat, forward-sloping surface when viewed in a second flow direction of the heat transfer fluid opposite the first flow direction. Clause 24. The power supply system according to clause 23, wherein the heat exchanger plate is: a first manifold channel coupled to the first heat exchanger port; a second manifold channel coupled to the second heat exchanger port; a plurality of cooling channels coupled between the first manifold channel and the second manifold channel and arranged to allow the heat transfer fluid to flow through the cooling channels. Clause 25. The power supply system of clause 24, wherein each cooling channel comprises the plurality of dimples configured to provide the turbulent flow of the heat transfer fluid through the cooling channel. Clause 26. A power supply system as described in clause 24 or 25, wherein the heat exchanger plate is configured to organize the plurality of cooling channels into a plurality of cooling units, and the plurality of cooling units are coupled in parallel between the first manifold channel and the second manifold channel. Clause 27. The power supply system of any one of clauses 24 to 26, wherein each cooling channel comprises a plurality of dimples arranged in a staggered pattern such that adjacent dimples in the first flow direction or the second flow direction extend from opposite sides of the cooling channel. Clause 28. The power supply system of clause 27, wherein adjacent dimples in the first flow direction or the second flow direction are recessed from laterally opposite sides of a top surface or a bottom surface of the cooling channel. Clause 29. The power system of clause 27, wherein the adjacent dimples in the first flow direction or the second flow direction are recessed from laterally opposite sides of the cooling channel. Clause 30. A power supply system according to any one of clauses 24 to 29, wherein the heat exchanger plate is configured to receive the heat transfer fluid from the first heat exchanger inlet / outlet and discharge the heat transfer fluid to the second heat exchanger inlet / outlet during a first operating period, and to receive the heat transfer fluid from the second heat exchanger inlet / outlet and discharge the heat transfer fluid to the first heat exchanger inlet / outlet during a second operating period before or after the first operating period. Clause 31. The power supply system according to any one of clauses 24 to 30, wherein the height of the first manifold channel and the second manifold channel is greater than the height of the cooling channel. Clause 32. The power supply system of any one of clauses 23 to 31, wherein the heat exchanger plate comprises aluminum. Clause 33. A method for cooling a battery, comprising: circulating a heat transfer fluid through a plurality of heat exchanger assemblies corresponding to a plurality of battery packs included in the electric aircraft by a fluid transport assembly, each heat exchanger assembly including a first heat exchanger inlet and a second heat exchanger inlet configured to receive or discharge the heat transfer fluid, and a heat exchanger plate; providing turbulent flow of the heat transfer fluid through a cooling channel of the heat exchanger plate with a plurality of dimples, the plurality of dimples in the cooling channel being staggered such that adjacent dimples in a flow direction of the cooling channel extend from opposite sides of the cooling channel; The operation of circulating the heat transfer fluid includes: receiving the heat transfer fluid from the first heat exchanger port and discharging the heat transfer fluid to the second heat exchanger port during a first operating period; and receiving the heat transfer fluid from the second heat exchanger inlet and discharging the heat transfer fluid to the first heat exchanger inlet during a second operating period. Clause 34. The method of clause 33, wherein the adjacent dimples in the flow direction are recessed from laterally opposite sides of a top or bottom surface of the cooling channel. Clause 35. The method of clause 33, wherein the adjacent dimples in the flow direction are recessed from laterally opposite sides of the cooling channel. Clause 36. The method of any one of clauses 33 to 35, wherein each of the plurality of dimples includes a symmetrical contour having a first substantially flat, forward-sloping surface when viewed in a first flow direction of the heat transfer fluid and a second substantially flat, forward-sloping surface when viewed in a second flow direction of the heat transfer fluid opposite the first flow direction. Clause 37. The method according to any one of clauses 33 to 36, The method further comprising balancing the heat transfer fluid entering the heat exchanger assembly with a plurality of flow restrictors in the fluid delivery assembly. Clause 38. The method of clause 37, wherein balancing the heat transfer fluid entering the heat exchanger assembly comprises: and ensuring substantially equal inlet pressures at each heat exchanger assembly by using sizing orifices for uniform flow distribution of the heat transfer fluid. Clause 39. The method of any one of clauses 33 to 38, further comprising balancing the heat transfer fluid entering the heat exchanger assembly through a plurality of differently sized conduit segments of the fluid transport assembly. Clause 40. The method of any one of clauses 33-39, wherein the heat transfer fluid comprises ethylene glycol water (EGW), ethylene glycol, polyethylene glycol, water, or any combination thereof. Clause 41. The method according to any one of clauses 33 to 40, wherein the dimples comprise a triangular or trapezoidal shape. Article 42. Aircraft: a plurality of battery packs installed in wings of the aircraft, each battery pack connected to an electric engine of the aircraft to provide power; a plurality of heat exchanger assemblies respectively connected to battery pack enclosures of the plurality of battery packs; a fluid delivery assembly coupled to the plurality of heat exchanger assemblies and configured to circulate a heat transfer fluid through the heat exchanger assemblies; each heat exchanger assembly comprising a first heat exchanger port and a second heat exchanger port configured to receive or discharge the heat transfer fluid, and a heat exchanger plate comprising a plurality of dimples configured to provide turbulent flow of the heat transfer fluid; each of the plurality of dimples includes a symmetrical profile having a first substantially flat, forward-swept surface when viewed from a first flow direction of the heat transfer fluid and a second substantially flat, forward-swept surface when viewed from a second flow direction of the heat transfer fluid opposite the first flow direction. Article 43. Power supply systems for aircraft, a plurality of battery packs configured to be installed within a wing of the aircraft, each battery pack configured to be connected to an electric engine of the aircraft to provide electrical power; a plurality of heat exchanger assemblies corresponding to the plurality of battery packs included in the electric aircraft, each heat exchanger assembly including a first heat exchanger inlet and a second heat exchanger inlet configured to receive a heat transfer fluid or discharge the heat transfer fluid; a fluid transport assembly coupled to the heat exchanger assembly, the fluid transport assembly including a first main inlet and a second main inlet, and configured to circulate the heat transfer fluid in parallel with the heat exchanger assembly, the first main inlet and the second main inlet being disposed at the same position on the electric aircraft, the fluid transport assembly comprising: The power system further comprising a plurality of conduit segments configured to balance the heat transfer fluid entering each of the heat exchanger assemblies. Article 44. A power supply system for an aircraft, comprising: a plurality of battery packs configured to be installed within a wing of the aircraft, each battery pack configured to be connected to an electric engine of the aircraft to provide electrical power; a plurality of heat exchanger assemblies respectively connected to battery pack enclosures of the plurality of battery packs, each heat exchanger assembly including a cooling channel; a fluid delivery assembly coupled to the plurality of heat exchanger assemblies and configured to circulate a heat transfer fluid through the cooling channels in the heat exchanger assemblies; 1. A power supply system, wherein each cooling channel comprises a plurality of dimples arranged in a staggered pattern such that adjacent dimples in a flow direction of the cooling channel extend from opposite sides of the cooling channel. Clause 45. The power system of clause 44, wherein the adjacent dimples in the flow direction are recessed from laterally opposite sides of a top or bottom surface of the cooling channel. Clause 46. The power system of clause 44, wherein the adjacent dimples in the flow direction are recessed from laterally opposite sides of the cooling channel. Clause 47. The power supply system according to any one of clauses 44 to 46, wherein each of the plurality of dimples includes a symmetrical contour having a substantially flat, forward-inclined surface when viewed from the flow direction. Clause 48. The power supply system of clause 47, wherein the dimple comprises a second substantially flat, forward-sloping surface when viewed in a direction opposite the flow direction. Clause 49. The power supply system according to any one of clauses 44 to 48, wherein the dimples include a triangular or trapezoidal shape. Article 50. Aircraft: An aircraft including a power supply system according to any one of clauses 44 to 49. Article 51. Aircraft: 23. An aircraft comprising a cooling system according to any one of clauses 1 to 22. Article 52. Aircraft: An aircraft including a power supply system according to any one of clauses 23 to 32. Article 53. Aircraft: The power supply system according to any one of clauses 44 to 49; A cooling system according to any one of clauses 1 to 22; and a power supply system according to any one of clauses 23 to 32.

[0165] The foregoing description has been presented for purposes of illustration. It is not exhaustive and is not intended to limit the disclosure to the precise forms or embodiments disclosed. Modifications and adaptations of the present disclosure will be apparent to those skilled in the art from consideration of the specification and practice of the disclosed embodiments of the disclosure disclosed herein.

Claims

1. 1. A cooling system for an electric aircraft, comprising: a plurality of heat exchanger assemblies corresponding to a plurality of battery packs included in the electric aircraft, each heat exchanger assembly including a first heat exchanger inlet and a second heat exchanger inlet configured to receive a heat transfer fluid or discharge the heat transfer fluid; a fluid transport assembly coupled to the heat exchanger assembly, the fluid transport assembly including a first main inlet and a second main inlet, and configured to circulate the heat transfer fluid in parallel with the heat exchanger assembly, the first main inlet and the second main inlet being disposed at the same position on the electric aircraft, the fluid transport assembly comprising: The cooling system further comprises a plurality of conduit segments configured to balance the heat transfer fluid entering each of the heat exchanger assemblies.

2. The cooling system of claim 1 , wherein the plurality of conduit segments comprises a plurality of flow restrictors configured to balance the heat transfer fluid entering each of the heat exchanger assemblies.

3. 3. The cooling system of claim 2, wherein each of the plurality of flow restrictors comprises a sizing orifice configured to ensure substantially equal inlet pressures at each heat exchanger assembly to achieve uniform flow distribution of the heat transfer fluid.

4. 4. The cooling system of claim 1, wherein each of the plurality of conduit segments includes a different inner diameter configured to balance the heat transfer fluid flowing into each of the plurality of heat exchanger assemblies.

5. 5. A cooling system according to any one of claims 1 to 4, wherein each heat exchanger assembly comprises:

1. A heat exchanger plate comprising: a first manifold channel coupled to the first heat exchanger port; a second manifold channel coupled to the second heat exchanger inlet / outlet.

6. 6. The cooling system of claim 5, a plurality of cooling channels coupled between the first manifold channel and the second manifold channel, the plurality of cooling channels arranged to allow the heat transfer fluid to flow through the cooling channels, each cooling channel comprising a plurality of dimples configured to provide turbulent flow of the heat transfer fluid through the cooling channel, each of the plurality of dimples comprising a symmetrical contour.

7. 7. The cooling system of claim 5 or 6, wherein the symmetrical contour comprises a first substantially flat, forward-inclined surface when viewed from a first flow direction of the heat transfer fluid and a second substantially flat, forward-inclined surface when viewed from a second flow direction of the heat transfer fluid opposite the first flow direction.

8. 8. The cooling system of claim 5, wherein the symmetrical contour comprises one of a triangle and a trapezoid.

9. 9. The cooling system of claim 5, wherein the heat exchanger plate is configured to organize the plurality of cooling channels into a plurality of cooling units, the plurality of cooling units being coupled in parallel between the first manifold channel and the second manifold channel.

10. 10. The cooling system according to claim 5, wherein the heat exchanger plate is configured to receive the heat transfer fluid from the first heat exchanger inlet / outlet and discharge the heat transfer fluid to the second heat exchanger inlet / outlet during a first operating period, and to receive the heat transfer fluid from the second heat exchanger inlet / outlet and discharge the heat transfer fluid to the first heat exchanger inlet / outlet during a second operating period before or after the first operating period.

11. 11. The cooling system of any one of claims 5 to 10, wherein the heat exchanger plate comprises aluminum.

12. 12. The cooling system of claim 5, wherein the plurality of dimples in each cooling channel are staggered such that adjacent dimples in the flow direction extend from opposite sides of the cooling channel.

13. The cooling system of claim 12 , wherein adjacent dimples in the flow direction are recessed from laterally opposite sides of a top or bottom surface of the cooling channel.

14. The cooling system of claim 12 , wherein the adjacent dimples in the flow direction are recessed from laterally opposite sides of the cooling channel.

15. 15. The cooling system of any one of claims 1 to 14, wherein the heat transfer fluid comprises ethylene glycol water (EGW), ethylene glycol, polyethylene glycol, water, or any combination thereof.

16. 16. The cooling system of claim 1, wherein the plurality of conduit segments of the fluid transport assembly comprises: a first tube and a plurality of first hoses coupled to the first tube; The cooling system further comprises a second tube and a plurality of second hoses coupled to the second tube.

17. 17. The cooling system of claim 16, wherein the first tube comprises a first trunk leading from the first main inlet / outlet, the plurality of first hoses comprise first branches coupled to the first trunk, the second tube comprises a second trunk leading from the second main inlet / outlet, and the plurality of second hoses comprise second branches coupled to the second trunk.

18. 20. The cooling system of claim 17, wherein a plurality of flow restrictors are coupled to the first hose or the second hose.

19. 19. The cooling system of any one of claims 16 to 18, wherein the first tube and the second tube comprise aluminum.

20. The cooling system according to any one of claims 1 to 19, wherein the first main inlet and the second main inlet are located at a top side of the electric aircraft.

21. 21. The cooling system of any one of claims 1 to 20, wherein the fluid transport assembly comprises: a first trunk path leading from the first main entrance to a first branch and a third branch; a second trunk path leading from the second main entrance to a second branch and a fourth branch; The cooling system, wherein the first trunk and the second trunk are substantially equal in length.

22. 22. The cooling system of claim 21, wherein the first main inlet and the second main inlet are located closer to the first branch than to the second branch.

23. 1. A power supply system for an aircraft, comprising: a plurality of battery packs configured to be installed within a wing of the aircraft, each battery pack configured to be connected to an electric engine of the aircraft to provide electrical power; a plurality of heat exchanger assemblies respectively connected to battery pack enclosures of the plurality of battery packs; a fluid delivery assembly coupled to the plurality of heat exchanger assemblies and configured to circulate a heat transfer fluid through the heat exchanger assemblies; each heat exchanger assembly comprising a first heat exchanger port and a second heat exchanger port configured to receive or discharge the heat transfer fluid, and a heat exchanger plate comprising a plurality of dimples configured to provide turbulent flow of the heat transfer fluid; each of the plurality of dimples includes a symmetrical profile having a first substantially flat, forward-sloping surface when viewed from a first flow direction of the heat transfer fluid and a second substantially flat, forward-sloping surface when viewed from a second flow direction of the heat transfer fluid opposite the first flow direction.

24. 24. The power system of claim 23, wherein the heat exchanger plate comprises: a first manifold channel coupled to the first heat exchanger port; a second manifold channel coupled to the second heat exchanger port; a plurality of cooling channels coupled between the first manifold channel and the second manifold channel and arranged to allow the heat transfer fluid to flow through the cooling channels.

25. 25. The power system of claim 24, wherein each cooling channel comprises the plurality of dimples configured to provide the turbulent flow of the heat transfer fluid through the cooling channel.

26. 26. The power supply system of claim 24 or 25, wherein the heat exchanger plate is configured to organize the plurality of cooling channels into a plurality of cooling units, the plurality of cooling units being coupled in parallel between the first manifold channel and the second manifold channel.

27. 27. The power supply system of claim 24, wherein each cooling channel comprises a plurality of dimples arranged in a staggered pattern such that adjacent dimples in the first flow direction or the second flow direction extend from opposite sides of the cooling channel.

28. 28. The power system of claim 27, wherein the adjacent dimples in the first flow direction or the second flow direction are recessed from laterally opposite sides of a top or bottom surface of the cooling channel.

29. 28. The power system of claim 27, wherein the adjacent dimples in the first flow direction or the second flow direction are recessed from laterally opposite sides of the cooling channel.

30. 30. The power supply system according to any one of claims 24 to 29, wherein the heat exchanger plate is configured to receive the heat transfer fluid from the first heat exchanger inlet / outlet and discharge the heat transfer fluid to the second heat exchanger inlet / outlet during a first operating period, and to receive the heat transfer fluid from the second heat exchanger inlet / outlet and discharge the heat transfer fluid to the first heat exchanger inlet / outlet during a second operating period before or after the first operating period.

31. 31. The power supply system according to claim 24, wherein a height of the first manifold channel and the second manifold channel is greater than a height of the cooling channel.

32. 32. The power system of any one of claims 23 to 31, wherein the heat exchanger plate comprises aluminum.

33. 1. A method for cooling a battery, comprising: circulating a heat transfer fluid through a plurality of heat exchanger assemblies corresponding to a plurality of battery packs included in the electric aircraft by a fluid transport assembly, each heat exchanger assembly including a first heat exchanger inlet and a second heat exchanger inlet configured to receive or discharge the heat transfer fluid, and a heat exchanger plate; providing turbulent flow of the heat transfer fluid through a cooling channel of the heat exchanger plate with a plurality of dimples, the plurality of dimples in the cooling channel being staggered such that adjacent dimples in a flow direction of the cooling channel extend from opposite sides of the cooling channel; The operation of circulating the heat transfer fluid includes: receiving the heat transfer fluid from the first heat exchanger port and discharging the heat transfer fluid to the second heat exchanger port during a first operating period; and receiving the heat transfer fluid from the second heat exchanger inlet and discharging the heat transfer fluid to the first heat exchanger inlet during a second operating period.

34. 34. The method of claim 33, wherein the adjacent dimples in the flow direction are recessed from laterally opposite sides of a top or bottom surface of the cooling channel.

35. 34. The method of claim 33, wherein the adjacent dimples in the flow direction are recessed from laterally opposite sides of the cooling channel.

36. 36. The method of any one of claims 33-35, wherein each of the plurality of dimples includes a symmetrical profile having a first substantially flat, forward-sloping surface when viewed in a first flow direction of the heat transfer fluid and a second substantially flat, forward-sloping surface when viewed in a second flow direction of the heat transfer fluid opposite the first flow direction.

37. The method according to any one of claims 33 to 36, The method further comprising balancing the heat transfer fluid entering the heat exchanger assembly with a plurality of flow restrictors in the fluid delivery assembly.

38. 38. The method of claim 37, wherein balancing the heat transfer fluid entering the heat exchanger assembly comprises: and ensuring substantially equal inlet pressures at each heat exchanger assembly by using sizing orifices for uniform flow distribution of the heat transfer fluid.

39. 39. The method of any one of claims 33 to 38, further comprising balancing the heat transfer fluid entering the heat exchanger assembly through a plurality of different sized conduit segments of the fluid transport assembly.

40. 40. The method of any one of claims 33 to 39, wherein the heat transfer fluid comprises ethylene glycol water (EGW), ethylene glycol, polyethylene glycol, water, or any combination thereof.

41. The method of any one of claims 33 to 40, wherein the dimples comprise a triangular or trapezoidal shape.

42. An aircraft, a plurality of battery packs installed in wings of the aircraft, each battery pack connected to an electric engine of the aircraft to provide power; a plurality of heat exchanger assemblies respectively connected to battery pack enclosures of the plurality of battery packs; a fluid delivery assembly coupled to the plurality of heat exchanger assemblies and configured to circulate a heat transfer fluid through the heat exchanger assemblies; each heat exchanger assembly comprising a first heat exchanger port and a second heat exchanger port configured to receive or discharge the heat transfer fluid, and a heat exchanger plate comprising a plurality of dimples configured to provide turbulent flow of the heat transfer fluid; each of the plurality of dimples includes a symmetrical profile having a first substantially flat, forward-swept surface when viewed from a first flow direction of the heat transfer fluid and a second substantially flat, forward-swept surface when viewed from a second flow direction of the heat transfer fluid opposite the first flow direction.

43. 1. A power supply system for an aircraft, comprising: a plurality of battery packs configured to be installed within a wing of the aircraft, each battery pack configured to be connected to an electric engine of the aircraft to provide electrical power; a plurality of heat exchanger assemblies corresponding to the plurality of battery packs included in the electric aircraft, each heat exchanger assembly including a first heat exchanger inlet and a second heat exchanger inlet configured to receive a heat transfer fluid or discharge the heat transfer fluid; a fluid transport assembly coupled to the heat exchanger assembly, the fluid transport assembly including a first main inlet and a second main inlet, the fluid transport assembly configured to circulate the heat transfer fluid in parallel with the heat exchanger assembly, the first main inlet and the second main inlet being disposed at the same position on the electric aircraft, the fluid transport assembly comprising: The power system further comprising a plurality of conduit segments configured to balance the heat transfer fluid entering each of the heat exchanger assemblies.

44. 1. A power supply system for an aircraft, comprising: a plurality of battery packs configured to be installed within a wing of the aircraft, each battery pack configured to be connected to an electric engine of the aircraft to provide electrical power; a plurality of heat exchanger assemblies respectively connected to battery pack enclosures of the plurality of battery packs, each heat exchanger assembly including a cooling channel; a fluid delivery assembly coupled to the plurality of heat exchanger assemblies and configured to circulate a heat transfer fluid through the cooling channels in the heat exchanger assemblies; 1. A power supply system, wherein each cooling channel comprises a plurality of dimples arranged in a staggered pattern such that adjacent dimples in a flow direction of the cooling channel extend from opposite sides of the cooling channel.

45. 45. The power system of claim 44, wherein adjacent dimples in the flow direction are recessed from laterally opposite sides of a top or bottom surface of the cooling channel.

46. 45. The power system of claim 44, wherein said adjacent dimples in said flow direction are recessed from laterally opposite sides of said cooling channel.

47. 47. The power supply system according to claim 44, wherein each of the plurality of dimples includes a symmetrical contour having a substantially flat, forward-sloping surface when viewed from the flow direction.

48. 48. The power system of claim 47, wherein said dimple comprises a second substantially flat, forward-sloping surface when viewed opposite said flow direction.

49. 49. The power supply system according to any one of claims 44 to 48, wherein the dimples include a triangular or trapezoidal shape.

50. An aircraft, An aircraft including a power supply system according to any one of claims 44 to 49.

51. An aircraft, An aircraft comprising a cooling system according to any one of claims 1 to 22.

52. An aircraft, An aircraft including a power supply system according to any one of claims 23 to 32.

53. An aircraft, The power supply system according to any one of claims 44 to 49; A cooling system according to any one of claims 1 to 22; and a power supply system according to any one of claims 23 to 32.

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