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 the priority and benefit of U.S. Patent Application No. 18 / 460,463, filed on September 1, 2023, entitled "HEAT EXCHANGER ASSEMBLIES AND COOLING SYSTEMS FOR EVTOL AIRCRAFT" (Attorney Docket No. 16163.0042 - 00000), and further claims the priority and benefit of U.S. Provisional Patent Application No. 63 / 383,660, filed on November 14, 2022, entitled "SYSTEMS AND METHODS FOR IMPROVED BATTERY ASSEMBLIES FOR EVTOL AIRCRAFT" (Attorney Docket No. 16163.6005 - 00000). The entire contents of the above applications are hereby incorporated by reference for all purposes.

[0002] Technical Field This disclosure generally relates to the field of powered aircraft. More specifically, without limitation, this disclosure relates to innovations in aircraft using electric propulsion systems. Certain aspects of this disclosure generally relate to high - voltage power supply (HVPS) systems and battery assemblies used in aircraft. Other aspects of this disclosure generally relate to improvements in cooling distribution systems that may also be used in other types of vehicles but can provide particular advantages for aircraft.

Summary of the Invention

[0003] A cooling system is provided according to several embodiments of the present disclosure. The cooling system comprises a plurality of heat exchanger assemblies corresponding to a plurality of battery packs and a fluid transport assembly. Each heat exchanger assembly comprises a first heat exchanger inlet and a second heat exchanger inlet configured to receive or discharge heat transfer fluid. The fluid transport assembly is coupled to the heat exchanger assemblies and configured to circulate the heat transfer fluid parallel to the heat exchanger assemblies in a U-flow configuration where the inlets and outlets of the heat transfer fluid are on the same side. The fluid transport assembly comprises a plurality of flow limiters configured to equalize the heat transfer fluid flowing into the heat exchanger assemblies. The heat transfer fluid flows through the corresponding flow limiters before flowing into the corresponding heat exchanger assemblies of the battery packs.

[0004] A power supply system for an aircraft is provided according to some embodiments of the present disclosure. The power supply system comprises a plurality of battery packs installed in the wing section of an aircraft, each battery pack being connected to two electric engines of the aircraft to supply power; a plurality of heat exchanger assemblies, each connected to a battery pack enclosure of the plurality of battery packs; and a fluid transport 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 comprises a first heat exchanger inlet and a second heat exchanger inlet configured to receive or discharge a heat transfer fluid; and a heat exchanger plate having a plurality of triangular dimples configured to provide turbulence for the heat transfer fluid.

[0005] Methods for cooling batteries are provided according to some embodiments of the present disclosure. The method involves circulating a heat transfer fluid through a fluid transport assembly to a plurality of heat exchanger assemblies corresponding to a plurality of battery packs included in an electric aircraft, each heat exchanger assembly comprising a first heat exchanger inlet and a second heat exchanger inlet configured to receive or discharge a heat transfer fluid, and a heat exchanger plate, and providing turbulence of the heat transfer fluid through cooling channels in the heat exchanger plate by a plurality of triangular dimples. The operation of circulating the heat transfer fluid includes, in 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 discharging the heat transfer fluid to

[0006] Please understand that the general description above and the detailed description below are for illustrative and explanatory purposes only and do not limit the embodiments disclosed in the claims. [Brief explanation of the drawing]

[0007] [Figure 1A] This figure shows a tiltrotor aircraft consistent with some embodiments of the present disclosure.

[0008] [Figure 1B] This figure shows a tiltrotor aircraft consistent with some embodiments of the present disclosure.

[0009] [Figure 1C] Figure 1B shows one exemplary embodiment of the battery.

[0010] [Figure 2] This block diagram shows a battery pack consistent with some embodiments of the present disclosure.

[0011] [Figure 3]A block diagram showing aspects of a high-voltage junction box (HVJB) consistent with some embodiments of the present disclosure.

[0012] [Figure 4] A cross-sectional view of an aircraft wing with a battery pack installed inside, consistent with some embodiments of the present disclosure.

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

[0014] [Figure 5B] A diagram showing components and units within a battery pack consistent with some embodiments of the present disclosure.

[0015] [Figure 6] A diagram showing an exemplary HVJB consistent with some embodiments of the present disclosure.

[0016] [Figure 7A] Shows an exemplary circuit diagram of an HVJB consistent with some embodiments of the present disclosure.

[0017] [Figure 7B] An exemplary block diagram showing the activation of a pilot fuse consistent with some embodiments of the present disclosure.

[0018] [Figure 8] A diagram showing an exemplary architecture of a form cell holder consistent with some embodiments of the present disclosure.

[0019] [Figure 9] A top view of a form cell holder consistent with some embodiments of the present disclosure.

[0020] [Figure 10A]A diagram showing an exemplary architecture of a current collector assembly that conforms to some embodiments of the present disclosure.

[0021] [Figure 10B] A diagram showing a top view of the current collector assembly of FIG. 10A at both the pack level and the cell block level that conforms to some embodiments of the present disclosure.

[0022] [Figure 11] An exemplary circuit diagram showing the series and parallel arrangements of battery cells forming the battery pack architecture shown in FIGS. 10A and 10B.

[0023] [Figure 12] A diagram showing an exemplary ventilation flap assembly that conforms to some embodiments of the present disclosure.

[0024] [Figure 13A] A diagram showing an enlarged portion of the ventilation flap assembly of FIG. 12 that conforms to some embodiments of the present disclosure.

[0025] [Figure 13B] A diagram showing a cross-sectional view taken along line A-A of FIG. 13A that conforms to some embodiments of the present disclosure.

[0026] [Figure 14] A diagram showing a cross-sectional view of a battery pack enclosure including a headspace directly above the ventilation flap assembly that conforms to some embodiments of the present disclosure.

[0027] [Figure 15] An exemplary heat exchanger plate is shown that conforms to some embodiments of the present disclosure.

[0028] [Figure 16A] The design of battery cooling in the HV cooling distribution subsystem is shown that conforms to some embodiments of the present disclosure. [Figure 16B] This document describes a battery cooling design in an HV cooling distribution subsystem consistent with several embodiments of the present disclosure. [Figure 16C] This document describes a battery cooling design in an HV cooling distribution subsystem consistent with several embodiments of the present disclosure. [Figure 16D] This document describes a battery cooling design in an HV cooling distribution subsystem consistent with several embodiments of the present disclosure.

[0029] [Figure 17] Figure 16C shows an exemplary battery cooling system based on the refrigerant line architecture consistent with several embodiments of this disclosure.

[0030] [Figure 18] This document illustrates an exemplary home system consistent with several embodiments of this disclosure.

[0031] [Figure 19] Figure 18 shows an exemplary electronic switching system for an exemplary home system, consistent with some embodiments of the present disclosure.

[0032] [Figure 20] This 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] This document presents an exemplary design of a cooling system consistent with some embodiments of the present disclosure. [Figure 21B] This document presents an exemplary design of a cooling system consistent with some embodiments of the present disclosure. [Figure 21C] This document presents an exemplary design of a cooling system consistent with some embodiments of the present disclosure. [Figure 21D] This document presents an exemplary design of a cooling system consistent with some embodiments of the present disclosure. [Figure 21E]This document presents an exemplary design of a cooling system consistent with some embodiments of the present disclosure.

[0034] [Figure 22A] An exemplary heat exchanger assembly for battery cooling in an HV cooling distribution subsystem, consistent with some embodiments of the present disclosure, is shown.

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

[0036] [Figure 22C] An exemplary refrigerant flow path within a heat exchanger assembly, consistent with some embodiments of the present disclosure, is shown.

[0037] [Figure 23A] This figure shows an enlarged portion of the cooling channel within a heat exchanger assembly, consistent with some embodiments of the present disclosure.

[0038] [Figure 23B] This figure shows a cross-sectional view along line AA, consistent with some embodiments of the present disclosure.

[0039] [Figure 24A] This figure shows an example of a dimple arrangement consistent with some embodiments of the present disclosure. [Figure 24B] This figure shows an example of a dimple arrangement consistent with some embodiments of the present disclosure. [Modes for carrying out the invention]

[0040] The following embodiments describe components of tiltrotor aircraft primarily used in non-conventional aircraft. For example, a tiltrotor aircraft may be intended for frequent (e.g., more than 50 flights per business day) and short-duration flights (e.g., less than 100 miles per flight) over, entering, and leaving densely populated areas. The aircraft may be intended to carry 4 to 6 passengers or commuters who expect a low-noise and low-vibration experience. Therefore, it may be desirable that the aircraft components be configured and designed to withstand frequent use without wear, generate little heat and vibration, and that the aircraft have mechanisms to effectively control and manage the heat or vibration generated by the components. Furthermore, some of these aircraft may be intended to operate in close proximity to each other over congested metropolitan areas. Therefore, it may be desirable that the aircraft components be configured and designed to generate low noise levels inside and outside the aircraft and 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 a single point 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 in a relatively limited space (e.g., a vertical takeoff and landing field, parking lot, or roadway) compared to a conventional airport runway, while carrying approximately 4 to 6 passengers or commuters with 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 aircraft components.

[0042] The disclosed embodiments provide novel and improved configurations of aircraft components not found in conventional aircraft, and / or specific design criteria for components that differ from conventional aircraft components. Such alternative configurations and design criteria, coupled with addressing the shortcomings and challenges of conventional components, have given rise to the embodiments disclosed herein for various configurations and designs of tiltrotor aircraft components.

[0043] In some embodiments, a tiltrotor aircraft may be designed to enable both vertical and conventional takeoffs and landings, while allowing vertical, forward, and transition flight through a distributed electric propulsion system. Thrust is generated by supplying high-voltage power to electric engines of the distributed electric propulsion system, each electric engine of which 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 electric engines connected to an onboard power supply, which may include devices capable of storing energy, such as batteries or capacitors, or one or more systems for utilizing or generating electricity, such as fuel-driven generators or solar panel arrays. Some disclosed embodiments provide weight reduction and space saving of components within the aircraft, thereby improving the efficiency and performance of the aircraft. With regard to passenger transport safety, the disclosed embodiments implement novel and improved safety protocols and system redundancy in the event of failure, minimizing any single point of failure in the aircraft propulsion system. Some disclosed embodiments also provide novel and improved approaches to meet aviation and transport laws and regulations. For example, the Federal Aviation Administration enforces federal laws and regulations requiring safety components, such as firewalls, adjacent to engines that use oil or other flammable materials in excess of a threshold amount.

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

[0045] As described herein, the orientation and use of electric propulsion systems may vary throughout the operation of the aircraft. In some embodiments, during vertical takeoff and landing, the forward and rearward propulsion systems may provide vertical thrust during takeoff and landing. In the flight phase when the aircraft is in forward flight mode, the rearward propeller system's propellers may be retracted to a fixed position to minimize drag, while the forward propulsion system may provide horizontal thrust. The rearward electric propulsion system may be actively retracted under position monitoring. Some embodiments may include transitions from vertical 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 the primary vertical direction in vertical flight mode and the predominantly horizontal direction in forward flight mode. Additional embodiments may include a variable pitch mechanism which may change the alignment angle of the propeller hub assembly blades of the forward propeller system for operation during the hovering phase, cruising phase, and transition phase. Some embodiments may include a conventional takeoff and landing (CTOL) configuration in which the tilter provides horizontal thrust for takeoff, cruising, and landing by the wings. 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 uncontrollable fires, such as utilizing less than one quart of flammable fluid or another harmless quantity of flammable fluid contained in both tilt and lift engines, having no nominal ignition sources within the electric engine, having an engine overheat operating limit more than 50°C below the autoignition temperature of the flammable fluid, overheat detection and protection, overvoltage detection and protection, and overcurrent detection and protection. In some embodiments, the design features of the electric engine may result in it being determined that it is not located in a designated fire-protected area.

[0047] As disclosed herein, an electric engine may comprise an inverter and motor, or an inverter, gearbox and motor in various configurations such as the representative configurations described herein. For example, an electric engine may comprise an electric motor, gearbox and inverter, all sharing the same central axis. In addition, the central axis may be configured along the axis of an output shaft connected to an aircraft propeller. In such exemplary configurations, the motor, gearbox and inverter all share the output shaft as the central axis and are oriented in a circular manner with respect to the output shaft. Additional embodiments may include a motor, gearbox and inverter mounted sequentially as a single unit, or a configuration in which some components such as the motor and gearbox are mounted as a single unit, while other components such as the inverter are located elsewhere, and the electric engine is connected using a wiring system.

[0048] As described herein, the electric engines for aircraft described herein may comprise some or all of a motor, inverter, and gearbox. Various configurations may include an inverter and motor such that the motor's output shaft directly provides speed and torque to the propeller shaft. Additional embodiments of the electric engine may comprise a motor, inverter, and gearbox in which the motor's output may be transmitted through a gearbox connected to the propeller's output shaft, and a motor, inverter, and gearbox in which 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. In addition, each configuration or orientation of the electric engine disclosed herein may include cooling via air cooling, coolant, or a mixture of both.

[0049] For example, the configuration of an electric engine may include a motor and an inverter, with the motor located between the aircraft's propeller and the inverter. In addition, the motor may include a gearbox. Furthermore, the inverter may share the same central axis as the motor and be housed in a cantilevered housing away from the rear of the motor, and may be air-cooled. Such an inverter orientation is recognized as not being the optimal configuration in terms of the housing required to achieve such a cantilevered orientation. In addition, the motor in this configuration utilizing air cooling may include potting material and air fins to assist in cooling the motor, which may result in a further increase in the system's mass.

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

[0051] Embodiments of an electric motor may include a stator housing, a wound stator assembly, a rotor, various bearings, and any additional components to assist in transmitting the speed and torque generated by the motor to a 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 through some or all of the individual components of the engine, such as the inverter, gearbox, or motor, to assist in managing the heat present in the engine. Additional embodiments may include using an air-cooling method to cool the electric engine or using a mixture of coolant and air to manage the heat generated during the 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, for example, using air and liquid cooling in the inverter and gearbox, or any other combination of air and liquid cooling over the inverter, gearbox, and motor, or even a subset of those components.

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

[0054] By using oil to not only lubricate but also cool the electric engine instead of a separate refrigerant, additional oil is added to the system, but it is understood that the oil eliminates the need for conventional parts 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 have 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, the overall mass of the system may decrease as fewer heat exchangers are used and other parts may become unnecessary, and a more attractive drag profile may exist. Furthermore, by using one substance for engine lubrication and cooling, the efficiency of the system may be improved by the reduction in mass and the advantage of cooling the engine with a substance instead of relying on air cooling, which may cause problems throughout the engine.

[0055] Additional embodiments of the electric engine may include various components to monitor any flammable fluids and ensure their ingress into specific parts of the electric engine. Some embodiments may include an electric engine with a wet zone enclosure that may be defined by a gearbox, motor, and / or heat exchanger. In some embodiments, the electric engine may have up to 4 liters of air in the motor gearbox housing that is in contact with the engine oil. Embodiments of the motor gearbox housing may use a breather to equalize the internal and external pressures. Embodiments of the breather may include a breather that protrudes above nearby design features to prevent the ingress of external fluids. Additional embodiments may include a breather with a screen and a roundabout inlet path to prevent the ingress of external foreign matter. Embodiments may include a sight glass, present in both tilt and lift electric engines, 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, stator winding sets, inverter bulk capacitors, power modules, control panel power modules, control panel control processors, control panel monitor processors, internal hot spots, and various other locations across the engine. Embodiments may include overheating limits considering known fault temperatures and operating limits related to the autoignition temperature of fluids. Some embodiments may include a high-voltage power system which may have fuses on high-voltage battery terminals that may rapidly and irreversibly disconnect the engine's electrical connections to mitigate overcurrent events. This overcurrent protection may operate when the electric engine's current consumption is greater than the overcurrent operation. Thus, in some embodiments, a fault condition leading to an overcurrent may only lead to a transient overheating, arc, or spark fault. Some embodiments may include a fire threat characteristic test ignition source which may be selected as an ignition source more serious than a short circuit occurring in the electric engine and disconnected by the engine fuse. In some embodiments, the inverter detects AC overcurrents to isolate the error phase and / or continuously monitors the input DC voltage and applies protective action to keep the voltage below the overvoltage limit.

[0057] For the purposes of this disclosure, the term “substantially” should be understood to mean a large or significant degree, such as being largely or within typical design, machining, and / or manufacturing tolerances as understood by a person skilled in the art.

[0058] Refer to Figure 1A, which shows a tiltrotor aircraft 100 consistent with several embodiments of the present disclosure. As shown in Figure 1A, in some embodiments, the distributed electric propulsion system of the tiltrotor aircraft 100 may comprise twelve electric engines 110 that can be mounted on forward and aft booms of the aircraft 100's wings. The forward electric engines 110 may be tiltable during flight between a horizontal orientation position (e.g., for generating forward thrust) and a vertical orientation position (e.g., for generating vertical lift). The forward electric engines 110 may be clockwise or counterclockwise with respect to the propeller rotation direction. The aft electric engines 110 may be fixed in a vertical orientation position (e.g., for generating vertical lift) and may also be clockwise or counterclockwise with respect to the propeller rotation direction.

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

[0060] In some embodiments, for vertical take-off and landing (VTOL) missions, the forward and rear electric engines 110 may provide vertical thrust during take-off and landing. During the flight phase when the aircraft 100 is in forward flight mode, the propellers of the rear electric engines 110 may be retracted to a fixed position to minimize drag, while the forward electric engines 110 may provide horizontal thrust. The rear electric engines 110 may be actively retracted under position monitoring. Transitions from vertical to horizontal flight, and vice versa, may be performed via a tilt propeller subsystem. The tilt propeller subsystem may redirect thrust between the primary vertical direction during vertical flight mode and the predominantly horizontal direction during forward flight mode. A variable pitch mechanism may change the angle of convergence of the propeller hub assembly blades of the forward electric engines for operation during the hovering, transition, and cruising phases.

[0061] In some embodiments, during a conventional take-off and landing (CTOL) mission, the forward electric engine 110 may provide horizontal thrust for take-off, cruising, and landing using the wings. In some embodiments, the rear electric engine 110 may not be used to generate thrust during a CTOL mission, and the rear propeller may be retracted into a predetermined position.

[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, inverter, and gearbox. In some embodiments, the motor, inverter, and gearbox may be interfaced to share a central axis. In some embodiments, the torque generated by the motor may be sent away from the propeller of the propulsion system to the gearbox. In some embodiments, the gearbox may provide gear reduction and then send the torque back to the propeller via the main shaft and bearings located inside the motor. In some embodiments, the inverter may be mounted at 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 to use a coolant such as oil to lubricate the motor, inverter, and / or gearbox, while sharing a common heat exchanger. In some embodiments, the amount of oil used for lubrication and cooling of the electric engine may vary, including less than one quart, less than two quarts, less than three quarts, or any other measured amount of oil.

[0063] In some embodiments, the tilt propeller system may include linear or rotary actuators for changing the orientation of the propulsion system during operation. In some embodiments, the pitch of the propulsion system may vary 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 can orient the propulsion system. 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 part of the redundant configuration to be driven by the motors, inverters, and gearboxes of another part of the configuration. In some embodiments, a gearbox configuration may also allow the tilt propeller system to maintain the orientation of the propulsion system with the assistance of additional power provided by the system.

[0064] As shown in Figure 1A, aircraft 100 is comprised of a distributed electric propulsion system that enables vertical flight, forward flight, and transitions. Six forward electric engines 110 (numbered 1-6 from left to right) enable vertical takeoff and landing, transition flight, and full-wing flight through the tilting of variable-pitch propellers. Six rear electric engines 110 (numbered 7-12 from left to right) are equipped with fixed-pitch propellers that operate during vertical takeoff and landing and transitions, and are retracted to the minimum drag position during normal flight. The flight control unit is an integrated fly-by-wire system characterized by envelope protection and structural load limiting functions. Aircraft 100 is equipped with advanced cockpit avionics, flight management systems, and sensors necessary to support the 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 electric engines 110, and thereby converting the HV power into mechanical shaft power used to rotate a propeller. As previously stated, the aircraft 100 described herein may have a plurality of electric engines 110 mounted on booms forward and aft of the wings. The amount of thrust generated by each electric engine 110 may be controlled by torque commands transmitted from a flight control system (FCS) to each electric engine 110 via a digital communication interface. Embodiments may include forward electric engines 110 whose orientation or inclination can be changed. Additional embodiments include forward engines which may be clockwise (CW) or counterclockwise (CCW). The forward electric engine propulsion subsystem may consist of a multi-blade adjustable-pitch propeller and a variable-pitch subsystem.

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

[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 supply the necessary power and energy required by the aircraft's equipment and may include a monitoring unit that measures performance parameters, runs estimation algorithms, and provides active protection against conditions that adversely affect the performance, lifespan, and safety of the HVPS. The HV vent subsystem is configured to safely remove battery exhaust during abnormal operation to prevent structural degradation and minimize the risk of diffusion. The HV cooling distribution subsystem is configured to distribute temperature-controlled coolant to the battery pack subsystem to reach the necessary temperatures for safe charging and discharging operations. The HV charging subsystem may provide an interface between the aircraft 100 and an external charger and may be responsible for sequencing charging operations and continuously monitoring voltage and current limits during charging sessions.

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

[0069] Refer to Figure 1B, which shows a tiltrotor aircraft 195 consistent with several embodiments of the present disclosure. As shown in Figure 1B, in some embodiments the tiltrotor aircraft may have 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, and 164 may include electrically independent high-voltage buses 130, 132, and 134 that connect pairs of batteries outside of the set of six batteries 101-106 for the given pair of battery pack units 160, 162, and 164. Figure 1C shows one exemplary embodiment of batteries 101-106. In some embodiments, each set of batteries in a pair may power four symmetrical electric engines 170-173, 174-177, or 178-181, and thereby simultaneously power four electric propeller units (EPUs) (labeled 1-12 in Figure 1B). In some embodiments, electric engines 170-181 may correspond to 12 electric engines (labeled 1-12 in Figure 1A) as shown in Figure 1A.

[0070] Referring to Figure 1C, some embodiments may include a pyrotechnic fuse 190 located inside each of the batteries 101-106. In some embodiments, the pyrotechnic fuse may be used to prevent a single point of failure in the event of a short circuit or failure of any one battery in a pair of batteries by disconnecting the faulty battery after the pyrotechnic fuse has activated, which disconnects the connection between the faulty battery and the working part of the electrical system of the tiltrotor aircraft 195.

[0071] Referring to Figure 1B, in some embodiments, additional fuses may be placed on the HV bus connecting other components of each pair of battery pack units. For example, fuses 110-121 may be placed on the HV bus between each battery 101-106 and their respective electric engines 170-181 and EPUs (labeled 1-12 in Figure 1B). Each of the fuses 110-121 may be used to prevent a single point of failure in the event of a short circuit or failure of any one component of one of the electric engines 170-181 or the EPU (labeled 1-12 in Figure 1B) by disconnecting the faulty component after one of the fuses 110-121 has been activated, which functions to disconnect the faulty component from the operable part of the electrical system of the tiltrotor aircraft 195. In another example, charging fuses 140, 142, 144, 146, 148 and 150 may be placed on the HV bus between each battery 101-106 and the charger or charging system. Each of the charging fuses 140, 142, 144, 146, 148, and 150 may be used, for example, to further prevent a single point of failure in the event of a short circuit or failure of any one battery during charging by disconnecting the faulty battery from the charger or charging system after one or more of the fuses 140, 142, 144, 146, 148, and 150 have been blown, which disconnects the connection between the faulty battery and the operable part 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 aircraft's operation depending on various failure scenarios. In some embodiments, each set of a pair of batteries may simultaneously power four symmetrical electric engines.

[0073] For example, referring further to Figures 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, and thus power two EPUs (labeled 1 and 12 in Figure 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, and thus power two other EPUs (labeled 4 and 9 in Figure 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 crosslink bus 130 may enable the first battery 101 to function as a backup battery for the second battery 102 (not shown), thereby supplying power to the two electric engines 172, 173 and EPU (labeled 4 and 9 in Figure 1B) that were initially powered by the second battery 102 when the second battery 102 (not shown) fails. In some embodiments, the first HV crosslink bus 130 may enable the second battery 102 (not shown) to function as a backup battery for the first battery 101, thereby supplying power to the two electric engines 170, 171 and EPU (labeled 1 and 12 in Figure 1B) that were 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 operating batteries in the battery pack pair.In some embodiments, disconnection may be performed via pyrotechnic fuses 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 unit 160. In some embodiments, each battery 101, 102 of the first paired battery pack unit 160 may be electrically isolated from the rest of the first paired battery pack unit 160 and 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 from the above example, and further referring to Figures 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, and thus power two other EPUs (labeled 2 and 11 in Figure 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, and thus power two other EPUs (labeled 5 and 8 in Figure 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 crosslink bus 132 may allow the third battery 103 to function as a backup battery for the fourth battery 104, thereby supplying power to the two electric engines 176, 177 and EPU (labeled 5 and 8 in Figure 1B) that were initially powered by the fourth battery 104 when the fourth battery 104 fails. In some embodiments, the second HV crosslink bus 132 may allow the fourth battery 104 to function as a backup battery for the third battery 103, thereby supplying power to the two electric engines 174, 175 and EPU (labeled 2 and 11 in Figure 1B) that were 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 operating 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 one output of each 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, each battery 103, 104 of the second pair of battery pack units 162 may be electrically isolated from the rest of the pair of battery pack units 162 and any other pairs of battery pack units 160, 164, for example via charging fuses 144, 146, so that a failure may be an isolated event that does not propagate along any connected HV channel.

[0075] Continuing from the above example and further referring to Figures 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, and thus power two other EPUs (labeled 3 and 10 in Figure 1B), and the sixth battery 106 of the third pair of battery pack units 164 may directly power two other electric engines 180, 181, and thus power two other EPUs (labeled 6 and 7 in Figure 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 crosslink bus 134 may enable the fifth battery 105 to function as a backup battery for the sixth battery 106, thereby supplying power to the two electric engines 180, 181 and EPU (labeled 6 and 7 in Figure 1B) that were initially powered by the sixth battery 106 when the sixth battery 106 fails. In some embodiments, the third HV crosslink bus 134 may enable the sixth battery 106 to function as a backup battery for the fifth battery 105, thereby supplying power to the two electric engines 178, 179 and EPU (labeled 3 and 10 in Figure 1B) that were 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 operating 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 one output of each 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, 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 failure may be an isolated event that does not propagate along any connected HV channel.

[0076] Continuing to refer to Figure 1B, in some embodiments, as a result of the above configuration example, the three pairs of battery pack units (i.e., the first, second, and third pairs of battery pack units) 160, 162, and 164 may be configured so that each pair of battery pack units 160, 162, and 164 powers one of the twelve electric engines, thereby enabling optimization of the power available to operate the twelve electric engines (labeled 1 to 12 in Figure 1B). For example, in some embodiments, one of each of the three pairs of battery pack units 160, 162, and 164 may be configured so that each set of batteries in the pair powers any four of the twelve electric engines, for example, one of the front outer electric engines, one of the rear inner electric engines, one of the front inner electric engines, and one of the rear 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 can power any combination of four forward and / or rear electric engines and / or EPUs. Appendix 1 provides additional information relating to embodiments of high-voltage power supply (HVPS) systems.

[0077] In some embodiments, as a result of the exemplary design described above, the backup battery may power all four electric engines simultaneously, particularly as the ability of one battery in each pair of battery packs to function as a backup battery for a failed battery in that pair of battery packs (i.e., the backup battery may power two electric engines that are directly electrically connected to the operating battery, and two electric engines via an HV crosslink bus connecting the pairs of batteries in the pair of battery pack units). In some embodiments, the backup battery may also power all four electric engines simultaneously at a slightly lower power level to maintain a power level sufficient for proper flight control. As a result of such exemplary design, all electric engines continue to receive power, and consequently, all motors and propellers continue to operate even if the first battery in any or all of the first, second, and third pairs of battery pack units fails. As a further consequence of such exemplary designs, approximately 20% less battery power may be required compared to independent or common bus structures to achieve the same results (e.g., to reach sufficient energy to controllably fly the design mission and sufficient thrust to safely perform a hovering landing). In some embodiments, as a further consequence of such exemplary designs and the approximately 20% reduction in power requirements, improved energy efficiency may also be achieved, for example, by longer-lasting power supply capabilities, extended aircraft range, and reduced aircraft weight based on the ability to utilize low-power batteries.

[0078] In some embodiments, a battery pack unit may include one battery to power any two electric engines and EPUs, and it is understood that the aircraft may have six battery pack units. As another example, in some embodiments, a battery pack unit may have 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 total of six electric engines and EPUs of the aircraft. As yet another example, in some embodiments, a battery pack unit may have 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 total of eight 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. In some embodiments, it is further understood that any combination of battery pack units of different sizes may be used to power twelve electric engines and EPUs. Furthermore, it is understood that any combination of battery pack units of different sizes may be used to power any number of electric engines and EPUs of an aircraft, without limiting the total number of electric engines and EPUs.

[0079] Refer to Figure 2. Figure 2 is a block diagram of 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 supply 280.

[0080] In the embodiment shown in Figure 2, the battery pack 120 comprises an HV junction box (HVJB) 122 and a cell stack assembly 124. Each battery pack 120 includes an HV distribution unit 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 to control the pack contactors and pyrofuses to protect from fault conditions and ensure safe operation. The BMS may also execute algorithms to determine the state of the pack (e.g., charge state, health state, 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 load 210 receiving HV power may include electric engines (e.g., two corresponding electric engines 110 connected to the battery pack 120 shown in Figure 1A), tilt actuators, DC / DC converters, ECS, or crosslinks used to connect pack pairs.

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

[0082] Refer to Figure 3. Figure 3 is a more detailed block diagram 300 showing the units within the HVJB122, consistent with some embodiments of the present disclosure. As shown in Figure 3, in some embodiments, the HV distribution unit 310 within the HVJB122 may include an 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 BMS320 within the HVJB122 may include a cell management unit (CMU)322 for monitoring the voltage of each set of seven parallel cells (i.e., 1S-7P cell groups) connected in series within a 14S-7P cell block. The CMU may also be used to monitor the temperature of the 14S-7P cell block. The CMU322 acquires measurements from all cell groups in the battery pack 120 and transmits the measurements to a battery management unit (BMU)324 via an isolated serial peripheral interface (isoSPI) in a daisy-chain configuration. In some embodiments, the CMU322 does not have an active management or control mechanism for the cells in the cell block, but the CMU322 can perform passive cell balancing of the series cell block when instructed by the BMU324. The BMU system architecture may provide 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 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 battery status even when not mounted on the aircraft 100. By monitoring the 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, low temperature, loss of electrical insulation, short circuits, and overcurrents.

[0085] The BMU324's diagnostic capabilities enable fault detection and isolation through embedded testing (BIT). Fault indications and overall usage are logged into non-volatile memory (NVM) for data retention and prediction. Furthermore, the BMU324 calculates the charge state (SOC), health state (SOH), power state (SOP), energy state (SOE), and temperature state (SOT) of the battery pack 120. The BMU324 also controls and monitors bus precharging and provides contactor instructions. In some embodiments, the BMU324 may use two microcontrollers to perform state calculations and contactor and fuse control functions so that a subset of the BMU324 functionality is retained when a failure occurs in a single microcontroller.

[0086] The CCU330 within the charging port assembly 262 may interface with an external battery charger and communicate with BMU324 on the six mounted battery packs 120. In some embodiments, the CCU330 may be a single PCBA with one microcontroller that manages the overall power supply to each battery pack 120 during charging. As shown in Figure 2, the CCU330 may perform a handshake between the ground charging subsystem 274 and the BMU324, requesting the BMU324 to open and close the contactor 312 and providing active detection and protection functions for overvoltage protection. In some embodiments, the CCU330 is located outside the battery packs 120. The BMU324 within each battery pack 120 maintains complete control and monitoring of those battery packs 120 during the charging operation.

[0087] Refer to Figure 4, and Figures 5A and 5B. Figure 4 is a cross-sectional view of the wing 400 of the aircraft 100 of Figure 1A, in 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 healthy state in which 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 discharge battery waste upward. The HVJB 122 of the battery pack 120 is mounted in the battery pack enclosure of the cell stack assembly 124.

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

[0089] In some embodiments, the battery cells 530 within the battery pack 120 may be lithium-ion cells, but the disclosure is not limited thereto. For example, the battery cells 530 may be 4.5Ah capacity cells having a cylindrical form factor of 2170. The battery cells 530 can operate in a voltage range of about 2.5V to about 4.18V, and each cell has a lifetime initial (BOL) capacity of 4.5Ah. In some embodiments, each battery cell 530 includes a current interruption device (CID) that permanently opens the battery cell 530 in the event of an increase in internal pressure.

[0090] Figure 6 shows an exemplary HVJB122 consistent with several embodiments of the present disclosure. As discussed above, the HVJB122 is configured to distribute high-voltage stored electrical energy for propulsion and other system needs, and the HVJB122 may include precharge resistors and relays, HV busbars, shunt resistors, HV contactors, connectors, a combination of active and passive fuses for protection against overcurrent and short-circuit conditions, and components to disable the high-voltage output in emergencies, if deemed necessary. The components within the HVJB122 may be sealed together with the BMU PCBA and mounted to a battery pack enclosure using fasteners.

[0091] Figure 7A shows an exemplary circuit diagram 700A of the HVJB122 consistent with several embodiments of the present disclosure. As described above, the HVJB122 may be electrically connected to an HV load 210 to supply HV power. Specifically, HV power can be supplied 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 Figure 7A, the DC / DC converter 710 and power storage element BT1 are connected to each of the HV loads 210 via precharge resistors (multiple) (e.g., resistor R1) or current sensing resistors (multiple) (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.). Active and passive fuses may be placed within the HV battery pack 120 to safely shut 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 supply when circuit disconnection and isolation are required. The BMS320 is used to operate pyrofuses F2, F3, F4, and / or F6 when a short-circuit event occurs. In some embodiments, the BMS320 in the HVJB122 can protect from overcurrent and electrically isolate the entire battery pack 120 from the connected load by sending a command signal to the corresponding pyrofuse driver to operate the pyrofuse when a fault occurs. In some embodiments, the HVJB122 may further provide redundant active triggers configured to allow the pyrofuse driver to operate one or more pyrofuses if the BMS320 fails to enable the pyrofuse driver, which are described below with reference to the accompanying drawings.

[0094] Figure 7B is an exemplary block diagram 700B showing the activation of a pyrofuse consistent with several embodiments of the present disclosure. As shown in Figure 7B, in some embodiments, the voltage output from the battery cell stack assembly 124 may 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 a pyrofuse redundant trigger (PRT) board 730 located within the BMS 320 and HVJB 122. The PRT board 730 may include various analog components to form a plurality of control logic circuits 732 to provide redundant active triggers to the corresponding pyrofuses (e.g., pyrofuses F2, F3, and F6). In some embodiments, the battery cell stack assembly 124 may also be electrically connected to an external DC-DC converter 720 located outside the battery pack 120. The external DC-DC converter 720 may also be configured to convert the voltage output from the battery cell stack assembly 124 to provide a supply voltage to the BMS 320 and PRT board 730. In other words, in the embodiment shown in Figure 7B, a redundant power supply system including an internal DC-DC converter 710 and an external DC-DC converter 720 may be used to ensure power supply and improve system stability. 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 supply power to other devices and units in the battery pack 120, ensuring that the system continues to operate normally.

[0095] As shown in Figure 7B, multiple pyrofuse drivers 734 configured to operate corresponding pyrofuses (e.g., pyrofuses F2, F3, and F6) may be located on the PRT board 730, but the disclosure is not limited thereto. In other embodiments, the pyrofuse drivers 734 may also be located separately on another circuit board. When the BMS 320 detects an overcurrent fault based on a signal detected by a corresponding current sensing resistor (e.g., resistor R3, resistor R4, or resistor R6), the 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 a pack pair). Therefore, the corresponding pyrofuse driver 734 may, after receiving the pyrofuse control command signal, activate the corresponding pyrofuse to electrically isolate the battery pack 120 from the load.

[0096] In some embodiments, if the BMS320 fails to function when a fault occurs and fails to output a pyrofuse control command signal, the pyrofuse will not blow and the power path will not be interrupted. In such cases, a control logic circuit 732, consisting of different thresholds for triggering the pyrofuse, may act as a redundant trigger, outputting a corresponding command signal to the pyrofuse driver 734 to activate the corresponding pyrofuse. Thus, the PRT board 730 is completely independent of the BMS320 and is designed to trigger and blow the pyrofuse based on analog circuitry. The redundant trigger source can further enhance the safety level of the HV battery pack 120.

[0097] As shown in Figure 7B, in some embodiments, both the BMS320 and the control logic circuit 732 in the PRT board 730 may use the same current sensor (e.g., sensing resistors R3, R4, and R6) to detect a short circuit. The BMS320 may be the primary source for sending an activation signal to the pyrofuse, while the PRT board 730 may play a role in checking for safety events and taking action only if the BMS320 fails to activate the pyrofuse.

[0098] Refer to Figure 8. Figure 8 shows an exemplary architecture of a foam cell holder 520 consistent with several embodiments of the present disclosure. In some embodiments, the cell holder 520 may be a machined foam with an FR4 top sheet and embedded busbars. In some embodiments, the cell holder 520 may be a precisely manufactured rigid, flame-retardant closed-cell polyurethane foam structure. When selecting the material for the foam cell holder 520, various properties are considered, including, for example, mechanical stiffness (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 capability (e.g., the ability to cast into complex shapes with minimal waste). For example, various structural foams may meet the standards of the aerospace industry. The cell holder 520 may have resistance 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. Therefore, the form cell holder 520 positions the battery cell 530 and separates the battery cell 530 from adjacent cells to protect it from thermal runaway. The cell holder 520 may also include nested busbars for connecting parts of the battery pack 120.

[0099] Figure 9 is a top view of a foam cell holder 520 consistent with several 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 in five columns 910-950. Each of columns 920-940 contains 42 rows of 7 cells, for a total of 294 cells per column. The first column 910 and the last column 950 have one fewer row, each containing a total of 287 cells. Each row of 7 parallel cells can be referred to as a cell group 960. Each cell is insulated from adjacent cells within the same cell group or adjacent cell groups by foam with a nominal thickness of 1.5 mm, reducing conductive heat transfer and the possibility of cell sidewall damage.

[0100] In some embodiments, the foam cell holder 520 includes alignment features and holes for positioning the part relative to a manufacturing jig. Channels along the edge of the foam are used as mounting locations for separate aluminum busbars, which are bolted and welded to the current collection assembly 540.

[0101] Figure 10A shows an exemplary architecture of a current collector assembly 540 consistent with several embodiments of the present disclosure. In some embodiments, the current collector assembly 540 may be a flexible circuit current collector assembly in which integrated sensing components are bonded and laser-welded. In some embodiments, the sensing components may be directly integrated into a flexible printed circuit board that is part of the current collector assembly without requiring bonding or welding. As shown in Figure 10A, the current collector assembly 540 may comprise five columns of cells and components integrated on a single central cell holder 520. In some embodiments, the current collector assembly 540 may provide a single component that integrates the current collector and sensing lines in a stacked manner to implement a stacked busbar having an integrated sensing layer for sensing the voltage and / or temperature of a corresponding cell group 542 or a corresponding cell block 544. For example, in each column, voltage and temperature sensing lines 546 may be arranged in a stacked manner and configured to collect voltage traces at one end of the column.

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

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

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

[0105] In some embodiments, the insulated ventilation flap may be positioned on top of the welded current collector assembly 540. Refer to Figure 12, and Figures 13A and 13B. Figure 12 shows an exemplary ventilation flap assembly 550 consistent with some embodiments of the present disclosure. Figure 13A shows a magnified portion of the ventilation flap assembly 550 of Figure 12, consistent with some embodiments of the present disclosure. Figure 13B shows a cross-sectional view along line AA of Figure 13A, consistent with some embodiments of the present disclosure. In Figures 13A and 13B, the ventilation flap assembly 550 is shown upside down with respect to the installation direction.

[0106] Specifically, as shown in Figures 12 and 13A and 13B, the vented flap assembly 550 comprises two layers 552 and 554 that are flexible and can withstand high temperatures (e.g., temporary or sustained high temperatures). For example, in the event of thermal runaway, the temperature may reach approximately 800°C. The first layer 552 may be, but is not limited to, a flexible silicone foam layer having PSA used to secure the vented flap assembly 550 to the current collector assembly 540. The second layer 554 is configured to protect adjacent cell groups 542 from cell ejection and can 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 Figure 12, the vented flap assembly 550 can be realized using a variety of designs. For example, in some other embodiments, the vented flap assembly 550 may use relatively rigid, non-cantilevered elements to provide a structure that may rupture under pressure without stretching too much outward in the event of a thermal runaway. The ventilation flap assembly 550 may employ various shapes that allow for ventilation by enabling deformation of the cantilevered portion(s). For example, a honeycomb structure may be applied to the arrangement of battery cells, but this disclosure is not limited thereto.

[0107] Figures 13A and 13B show the series connections 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., about 20 mm in height) between the pack lid 570 and the second layer 554 to vent gas 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 venting gas from a cell, the gas vented from the affected cell is contained by a foam cell holder 520, which enters a space below the cell and is surrounded by a silicone foam layer 552 and a polyamide thermoplastic layer 554, and directed downward. This space may also be called 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 the pressure in the ventilation manifold 556 increases, flaps on the polyamide thermoplastic layer 554 open, allowing gas and potential ejecta to enter the headspace 558 below the ventilation flap assembly 550. Thus, the ventilation flap assembly 550 may provide a one-way valve positioned above the cell and current collector assembly 540 to safely ventilate a thermally runaway cell and protect adjacent cells from the heat and ejecta of the failed cell.

[0108] Figure 14 shows a cross-sectional view of the battery pack enclosure 1400, including headspace directly above the ventilation flap assembly 550, consistent with several embodiments of the present disclosure. As with Figures 13A and 13B, the structure in Figure 14 is shown upside down with respect to the installation direction. As shown in Figure 14, the battery pack enclosure 1400 provides a structural interface for mounting the battery pack 120 to the main structure of the aircraft, housing the entire 208S-7P cell stack assembly 124 and connecting to the 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] Therefore, the battery pack enclosure 1400 has a defined path for venting ventilation gases via a pressure relief safety device (e.g., a burst valve) that prevents overpressurization 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 surrounding environment during normal operation. The vent interfaces with a duct that discharges battery exhaust to the outside of the aircraft 100. Apart from the ventilation path, the battery pack enclosure 1400 may be sealed to prevent leakage of cell electrolyte from being contained within the enclosure and flowing into the aircraft 100, which could pose a hazard.

[0110] Figure 15 shows an exemplary heat exchanger plate 1500 consistent with several embodiments of the present disclosure. The heat exchanger plate 1500 may be fixed to a battery pack enclosure 510 to allow temperature control within the battery pack 120 using a thermoconditioned 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 control can be used only while the aircraft 100 is on the ground. The heat exchanger plate 1500 may be sized to support any required heating or cooling needs while operating under nominal pressure and flow conditions. For example, the heat exchanger plate 1500 can be bonded to the battery pack enclosure 510 using a thermally conductive adhesive (e.g., a flame-retardant acrylic adhesive). The thermally conductive adhesive can be used to conduct heat generated by the cells to the heat exchanger plate 1500.

[0111] Figures 16A to 16D show different designs of battery cooling in the HV cooling distribution subsystem, respectively, consistent with several embodiments of the present disclosure. Some item numbers are shown in Figure 16A and may be omitted in subsequent figures for clarity. As previously stated, the aircraft 100 may have, for example, six battery packs 120. In some embodiments, a different number of battery packs 120 may be provided, for example, two, four, eight, or ten battery packs 120. In some embodiments, an even number of battery packs may be provided to achieve a symmetrical distribution of weight 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 housed in a central position of the aircraft 100, a central battery pack 120 may be provided, in which the same number of battery packs 120 are arranged on either side of the central battery pack 120.

[0112] Each battery pack 120 may be provided with a heat exchanger inlet 1621 and a heat exchanger outlet 1622 for circulating a refrigerant to achieve battery cooling. Figures 16A to 16D show different refrigerant line architectures which may provide a single main inlet 1610 and a 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 to circulate it through the inlet path 1611 and the outlet path 1651. The inlet path 1611 and the outlet path 1651 may comprise conduits formed from, for example, rigid or semi-rigid piping, flexible hoses, integrally molded channels, or a combination of these or other suitable materials. The assembly of such piping and associated fixtures, fittings, inlets and outlets for distributing refrigerant to the heat exchangers within each battery pack 120 may be called a fluid transport assembly. The configuration according to the embodiments of this disclosure makes it possible to distribute refrigerant evenly and in parallel to multiple battery packs 120 installed inside the wings of aircraft 100 without requiring active control components. The arrangements in Figures 16A to 16D may be configured to reduce the overall diameter, length, and number of joints of the inlet path 1611 and outlet path 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 to and disconnection from a ground-based refrigerant supply source, thereby enabling a rapid turnaround between landing and takeoff. This may thus favorably reduce downtime in the flight schedule. Alternatively, the cooling time may be effectively extended during a given landing period. In some embodiments, battery cooling may be achieved without using an onboard pump or other active flow control components. Instead, an external pump (e.g., on a ground-based refrigerant supply system outside the aircraft 100) may be used to charge the battery packs 120 while simultaneously circulating the refrigerant within the battery packs 120.

[0113] In some embodiments, the cooling operation may be performed without exceeding the overall system pressure limit (e.g., less than 100 psi), thus eliminating the need for special connectors. The cooling system may also be a passive system without additional components for performing an active control method to equalize the flow across multiple battery packs 120. Thus, the weight of the system can be minimized. Furthermore, the passive cooling system may offer a simplified design, reducing manufacturing costs and eliminating potential points of failure.

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

[0115] In some embodiments, the main inlet 1610 and main outlet 1650 may be located adjacent to each other or within a single fixture so that the supply and return lines for a ground-based refrigerant supply may be mounted by a single nozzle attachment. Furthermore, in some embodiments, the main inlet 1610 and main outlet 1650 may be manufactured in conjunction with other maintenance attachments, such as a charging port, to enable quick connection of all maintenance utilities. This may enable simpler maintenance operations or allow operations to be performed by autonomous devices. Furthermore, while ground-based cooling supplies have been 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 rapid 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 supply or charge the aircraft 100. This may enable, for example, longer continuous flight or provide emergency maintenance services in areas where landing is not optimal or impossible. Therefore, in some embodiments, the main inlet 1610 and main outlet 1650 may be provided within a self-docking fixture, along with a charging port, for attachment and detachment during flight.

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

[0117] In some embodiments, the U-shaped flow design may feature a "first-in, last-out" architecture such that the first battery pack 120 receiving refrigerant from the inlet path 1611 in the inlet flow direction may be the last battery pack 120 discharging refrigerant to the outlet path 1651 in the outlet flow direction. As a result, there will be differences in the path length of the individual refrigerant flow loops passing through each individual battery pack. For example, as seen in Figure 16A, the refrigerant flowing through the rightmost battery pack 120 may follow the shortest loop between the main inlet 1610 and the main outlet 1650, while the refrigerant flowing through the leftmost battery pack 120 may follow the longest loop. This U-shaped flow arrangement allows for a desired arrangement of the main inlet 1610 and main outlet 1650 in close proximity to each other, but can lead to pressure and flow imbalances if not properly considered. Some embodiments of this disclosure provide systems and methods for passively maintaining a balanced flow while keeping the main inlet 1610 and main outlet 1650 integrated.

[0118] Figure 16B shows an alternative “Z-flow” type 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 on the other side, forming a roughly “Z” shape in the schematic diagram. The Z-flow type refrigerant line architecture 1600B may also allow the refrigerant path length to be the same for each of the six battery packs 120, which may be desirable to optimize the 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 have a “first-in, first-out” structure, such that a first battery pack 120 that receives refrigerant from the inlet path 1611 in the inlet flow direction may be a first battery pack 120 that discharges refrigerant to the outlet path 1651 in the outlet flow direction along the refrigerant flow path. As a result, the total length of the refrigerant flow path is substantially the same 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-shaped flow refrigerant line architecture 1600A or the Z-shaped flow refrigerant line structure 1600B, the flow distribution within the battery packs 120 may be balanced by varying the inner diameter between different segments of the inlet or outlet paths, for example. For instance, the inlet path segment immediately upstream of the rightmost (first) battery pack 120 in Figure 16B may be sized to supply refrigerant to all battery packs 120, while the inlet path segment immediately upstream of the leftmost (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 decrease continuously in the direction of the inlet path. In some embodiments, the reverse arrangement may be configured for the outlet path in Figure 16B. For example, the inner diameter of the outlet path segments may increase continuously in the direction of the outlet path. Thus, a refrigerant flow architecture is provided that uses minimal piping or hose sizes in each segment along the supply and return paths, thereby further minimizing the wet and dry weight of the cooling system.

[0120] To maintain a balanced flow while keeping the main inlet 1610 and main outlet 1650 on the same side of the aircraft 100, an improved "U-flow" type cooling system 1600C is provided in Figure 16C, in which customized flow limiters 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. In other words, the refrigerant may flow through the customized flow limiters 1630 before flowing into the corresponding battery packs 120. Each customized flow limiter 1630 may provide desired flow resistance to achieve an appropriate amount of refrigerant flow and to balance the flow flowing into each of the battery packs 120. For example, in the cooling system 1600C, the flow limiters 1630 may include sizing orifices to equalize the pressure drop from the common main inlet 1610 to each battery pack 120. This may ensure a uniform flow distribution of the refrigerant to each battery pack 120. In this way, the cooling architecture may function efficiently without requiring multiple pipes of different diameters or equivalent path lengths.

[0121] Figure 16D shows an improved "Z-flow" type cooling system 1600D with an added in-flight return path 1640. The in-flight return path 1640 may include additional landing gear for an outlet path 1651, for example, located inside the wing of the aircraft 100. By providing the in-flight return path 1640, the main inlet 1610 and main outlet 1650 may be located on the same side of the aircraft 100, while the flow distribution within the battery pack 120 may still be balanced, for example by using the Z-flow architecture. Similar to the Z-flow type cooling system 1600B, in some embodiments this design may include piping segments of different sizes or multiple orifices of different sizes.

[0122] Additionally, by providing segments with continuously increasing / decreasing inner diameters, using flow limiters, and / or providing each loop with an equivalent path length as described above, embodiments of the present disclosure may realize a symmetric architecture configured to provide uniform flow characteristics whether the refrigerant is circulating in the forward or reverse direction. For example, in some embodiments, the refrigerant architecture may be configured to have symmetrical flow in both directions. This may allow for improved cooling uniformity by circulating the refrigerant alternately in different directions, as will be discussed further below with respect to Figure 20, for example.

[0123] Therefore, although the elements in Figures 16A to 16D and other figures may be referred to by terms such as “inlet” and “outlet,” it should be understood that in some embodiments, these elements may be configured to allow refrigerant to flow in either direction. For example, in some embodiments, the main inlet 1610 may function as a first main inlet / outlet, the main outlet 1650 may function as a second main inlet / outlet, the heat exchanger inlet 1621 may function as a first heat exchanger inlet / outlet, and the heat exchanger outlet 1622 may function as a second heat exchanger inlet / outlet. Furthermore, in some embodiments, the entire cooling system between the first and second main inlets / outlets is not provided with check valves or other unidirectional flow fixation devices to enable such reverse flow operation. Additionally, in some embodiments, to ensure symmetrical flow characteristics in the forward and reverse flow directions, the inlet path 2011 and the outlet path 1651 may have substantially equal lengths.

[0124] Figure 17 shows an exemplary battery cooling system 1700 based on the cooling system of Figure 16C, consistent with several embodiments of the present disclosure. As previously stated, the aircraft 100 may have a common main inlet 1710 and a common main outlet 1750 for receiving and discharging the refrigerant. The refrigerant may flow through their respective flow limiters 1730 added to the tubes associated with the multiple battery packs, and then flow in parallel into each of the multiple battery packs 120. In some embodiments, each flow limiter 1730 may have different lengths and / or orifice sizes depending on the desired flow characteristics. For example, the orifice size may increase along the inlet flow direction to compensate for the pressure drop that occurs as the refrigerant branches toward the upstream battery packs 120. As a non-limiting example for illustrative purposes, flow limiters 1730a-f may have orifice sizes in the inlet flow direction, for example, 3.7 mm, 4.1 mm, 4.65 mm, 6.8 mm, 10 mm, and 12.5 mm, respectively, and the tube diameter may be, for example, about 20 mm or 0.5 inches. The flow limiter may be located upstream or downstream of the battery pack 120. For example, an embodiment of the flow limiter 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, about 60 liters / minute (LPM). That is, for each of the six battery packs 120, the individual flow rate may be about 10 liters / minute. However, the disclosure is not limited thereto. 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 passing through the heat exchanger of each battery pack 120 may deviate from the average value of all battery packs 120 by, for example, 5%, 2%, 1%, or 0.5%. The cooling system may also be robust to errors in refrigerant flow and distribution. For example, continuing the above example, if the refrigerant is distributed among the six battery packs 120 at twice the flow rate of 120 liters / minute due to uneven distribution within the cooling system, each pack may accept an average flow rate of 20 liters / minute. For example, the differential pressure of the battery packs 120 may be about 20 psi when the flow rate is 10 liters / minute and about 40 psi when the flow rate is 20 liters / minute.

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

[0127] Refer to Figure 18. Figure 18 shows an exemplary home system 1800 consistent with several embodiments of the present disclosure. Batteries and / or battery systems used in demanding applications within aircraft, such as high-voltage batteries manufactured primarily for powering eVTOL aircraft, are understood to undergo progressive degradation of their battery characteristics, rendering them unusable for their primary application over time. Typical parameters indicating 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 operation, an increase in the internal resistance / impedance of the battery or battery system, and a decrease in power supply capacity below the requirements of the aircraft's host devices. However, such high-voltage batteries or battery systems may remain useful in other secondary applications, including, but not limited to, household power systems, rural energy storage systems, and other backup power systems, including large-scale power grids and household power grid backup systems. For example, secondary uses of such batteries and / or battery systems may include supplying power to a home, the reused batteries or battery systems may be used as an energy source to supply power to the home, and the batteries or battery systems may be directly connected to the inverters and power supplies of the home system.

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

[0129] In some embodiments, the secondary use may be installed within a home system (i.e., a household power system) and supply power to the home system. In some embodiments, the home system may be a utility system that supplies power to various subsystems of a house, housing complex, commercial facility, or industrial facility. In some embodiments, the home system may include an electronic switching system (ESS). As shown in the example in Figure 18, the home system 1800 may include a refurbished battery 1810, which is coupled to a battery interface 1815, which is coupled to an inverter 1820, which is coupled to one or more line filters 1825. In some embodiments, the battery 1810, battery interface 1815, inverter 1820, and one or more line filters 1825 may include an electronic switching system 1860 of the home system 1800. In some embodiments, the battery interface 1815 may be a DC / DC interface. In some embodiments, such exemplary electronic switching system 1860 may be connected to one or more household power 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 supplied to the household power lines 1840 via either the electronic switching system 1860 or an external power grid 1850.

[0130] Refer to Figure 19. Figure 19 shows an exemplary electronic switching system 1900 for the home system 1800 shown in Figure 18. As shown in the example in Figure 19, the electronic switching system 1900 may comprise a repurposed HV battery and a battery interface 1920 (the battery interface includes a battery management system (BMS)), and power electronics equipment 1910 including an inverter and one or more line filters. In some embodiments, such exemplary electronic switching system 1900 may be connected to one or more household power lines via a switching network (i.e., switchgear). In some embodiments, the switching network may be configured to supply power to the household power lines via either the electronic switching system or an external power grid. In some embodiments, the inverter in 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 Figures 18 and 19, in some embodiments, the refurbished battery may be a high-power (i.e., HV) battery equipped with high-power cells, which may enable increased power applications in the home electrical system 1840. As an example, one or more such HV batteries equipped with high-power cells may be used for DC fast charging (DCFC) applications in the home. In some embodiments, the DCFC application may include DCFC of an electric vehicle (EV) at a speed significantly greater than currently possible. For example, two high-power batteries equipped with the high-power cells described herein may be used in a home system to charge an EV at a speed about 25 times faster than an existing home EV charging station, in about 40–80% of the time typically required by a V3 supercharger. As another example, to further support the DCFC application via an electronic switching system 1900, a custom DC / DC battery interface may be provided to support DCFC performance as required. As a result, in some embodiments, the custom DC / DC battery interface may include control logic, sensors, and / or communication interfaces that are minimized in size, optimized in efficiency, and / or further support the vehicle's DCFC via, for example, a household electronic switching system 1900.

[0132] Referring further to Figure 19, in a further embodiment where the refurbished battery is a high-power battery with high-power cells, the refurbished battery may have even lower resistance. In some embodiments, the refurbished battery may generate even less heat compared to the battery in an existing household electronic switching system. As a result, in some embodiments, the refurbished battery may be used as a heat sink for the inverter. As a further consequence, in some embodiments, such an inverter may not require any additional forced cooling for operation. As a further consequence, 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] Continuing to refer to Figures 18 and 19, in some embodiments, the home system comprises a parallel battery 1810, at least one of which is a refurbished battery. In some embodiments, no additional equipment may be required to add further capacity to the electronic switching system 1860 by providing additional batteries connected in parallel to the first battery 1810 of the electronic switching system 1860. For example, in some embodiments, the electronic switching system 1860 may further comprise a master battery comprising 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, a specific capacity required for a particular home may be met by providing batteries connected in parallel to the master battery.

[0134] In some embodiments, the refurbished battery may have most of the sensors necessary to function properly within the electronic switching system 1900 and / or home system 1800. For example, a refurbished battery for use in the electronic switching system 1900 and / or home system 1800 may only require the replacement of the aircraft-specific high-voltage junction box (HVJB) with a home system-specific HVJB. In some embodiments, a refurbished battery for use in the 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 implemented in the battery from the initial construction for its primary use. In such embodiments, the common HVJB may eliminate the need to replace the HVJB with a home system-specific HVJB before the battery's secondary use. Because overload and / or storage capacity is required for aircraft batteries, such batteries may also have a capacity that can be preloaded to set up home system compatibility and / or functionality during their initial construction.

[0136] Figure 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 mentioned, the aircraft 100 may comprise, for example, six battery packs 120. Each battery pack 120 may have inlets and outlets for circulating a refrigerant within the heat exchanger assembly of the battery pack 120 to achieve battery cooling. In Figure 20, similar to the embodiments of Figures 16-17 described above, a refrigerant line architecture is proposed with refrigerant inlets and outlets to circulate the refrigerant in parallel to the six battery packs 120 installed in the wings of the aircraft 100 in order to minimize both the dry and wet weight of the cooling system.

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

[0138] For example, a dual U-flow system may include a first main inlet / outlet 2010 connected to a first trunk 2013 and a second main inlet / outlet 2050 connected to a second trunk 2053. The first trunk and the second trunk may each branch, for example, at the central position of the aircraft 100, into a first U-flow loop providing a first set of battery packs 120 in the first wing of the aircraft 100 and a second U-flow loop providing a second set of battery packs 120 in the 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, in order to ensure symmetrical characteristics in the forward and reverse flow directions, the first trunk 2013 and the second trunk 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 main 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 main 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 to 17, the cooling system 2000 may be configured as a symmetrical system that can circulate the refrigerant flow in either direction, so that any set of trunks and branches may be used as either inlet or outlet lines.

[0140] Providing refrigerant flow in alternating directions during a continuous circulation period may improve cooling efficiency or uniformity. For example, in a given cooling operation, the flow direction may be reversed once or multiple times. For example, the flow direction may be changed in the middle of a cooling cycle, or at other periodic or irregular intervals. In some embodiments, the refrigerant may flow mainly in one direction, with periodic interruptions by reciprocating flow. An external refrigerant supply source may be configured to control such changes in flow direction, for example, by the operation of a pump, a flow-reversal valve, etc.

[0141] The first and second main inlets of the transport assembly 2001 may be integrally located at the same position on the aircraft 100. For example, the first and second main inlets may be located on the top, bottom, or side of the fuselage of the aircraft 100, or within the wings. For example, locating them on the top of the fuselage, for example between or behind the wings, may favorably reduce the overall length of the hoses because they are closest to the battery pack 120. However, locating them on the side or bottom of the fuselage may allow for easy connection access. In some embodiments, if the first and second main inlets are located in asymmetrical positions, such as on one side of the fuselage, the inlet and outlet channels 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 channels may terminate at their respective branches at a central position on the aircraft 100. In this way, symmetrical flow path lengths may be maintained while providing an asymmetric connection to an external coolant supply source. As described above, the first main inlet and the second main inlet may be configured as a single connection port, so that an external refrigerant supply source may simultaneously provide a flow of equilibrium refrigerant to all battery packs of the aircraft 100 through a single connection.

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

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

[0144] Figures 21A and 21B show example designs of a cooling system 2100 consistent with several embodiments of the present disclosure. The cooling system 2100 may correspond, for example, to the cooling system 2000 in Figure 20. As shown in Figure 21A, the fluid transport assembly 2101 may provide battery coolant to each battery pack 120, respectively. The fluid transport 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 be divided into a first branch, a second branch, a third branch, and a fourth branch, respectively, as described above. As shown in Figure 21A, the first and second main inlets of the fluid transport assembly 2101 may be located, for example, at the top of the fuselage in the center of the wingspan of the aircraft 100, to achieve a more balanced cooling distribution with three battery packs 120 on one side and three battery packs 120 on the other side. Alternatively or additionally, as described above, the first and second main inlets may be located, for example, on one side of the fuselage or on one wing.

[0145] Figure 21B shows a further illustration of the cooling system 2100 according to an embodiment of the present disclosure. The illustration in Figure 21B can be considered an inverted view of Figure 21A, showing the lower side of the cooling system 2101. Figure 21B shows details of the various conduit sections of the fluid transport system 2101 and their connections 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 limiter (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 transport assembly 2101 may be made of various materials. For example, as shown in Figure 21B, the first main channel 2113 and the second main channel 2152 may include rigid tubing material, such as aluminum or other high-strength, lightweight material. Additionally, the first segment 2112a of the first branch 2114 may be made of the same or similar material. Subsequent segments 2112 may be made of different materials, such as flexible hose. Similarly, the second segment 2152a of the second branch 2154 may be made of the same or similar material. Subsequent segments 2152 may be made of different materials, such as flexible hose. Configuring the downstream segments of the branches in the fluid conveying 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 with the body of the aircraft 100. Instead of hoses, polyethylene or other semi-rigid tubing may be provided. In some embodiments, all conduits in the fluid conveying assembly may be made of the same material.

[0146] Figure 22A shows an exemplary heat exchanger assembly 2200 for battery cooling in an HV cooling distribution subsystem, consistent with several embodiments of the present disclosure. Figure 22B is a top view of the heat exchanger assembly 2200 shown in Figure 22A, consistent with several embodiments of the present disclosure. The design of the heat exchanger assembly 2200 shown in Figures 22A and 22B allows for cooling from the bottom of multiple battery cells within a 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 embodiments of Figures 22A and 22B, the heat exchanger assembly 2200 may be integrated into a thermal management system to ensure maximum cooling capacity without exceeding a desired overall system pressure limit (e.g., less than 100 psi), and may not require special or customized connectors. Furthermore, by minimizing the thermal gradient between cells, this design may also extend the battery life of the battery pack 120.

[0147] In particular, in the heat exchanger assembly 2200, multiple dimples 2210 are incorporated onto the heat exchanger plate (i.e., cooling plate) 2220 to function as turbulence generators in order to improve the heat transfer efficiency of the heat exchanger assembly 2200. For example, the dimples 2210 have formed depressions within the cooling channels 2222 of the heat exchanger assembly 2200, configured to generate calculated strain within the internal volume to disrupt the laminar flow of refrigerant, and may more uniformly distribute surface contact between refrigerant 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 generated, for example, by a hydroforming process or subsequently engraved. The heat exchanger plate 2220 may be, for example, an aluminum plate or another lightweight, strong, and thermally conductive material suitable for cost-effective manufacturing. This approach enables 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 inlet / outlet 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 inlet / outlet 2250 for receiving or discharging a 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 so 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 turbulence for the heat transfer fluid through the cooling channel 2222. The arrangement of the dimples 2210 may be configured to meet the specific needs of, for example, an electric aircraft. For example, the dimples 2210 may be symmetrical in both shape and arrangement along the flow direction, so that their heat transfer characteristics are identical in either flow direction. This may improve the performance of the bidirectional cooling operation described above.

[0150] Furthermore, the dimples 2210 may be configured to introduce local turbulence into the refrigerant flow without excessively disturbing the laminar flow. For example, if the dimples 2210 excessively separate the refrigerant flow on the rear side of the dimples 2210, the overall cooling efficiency and uniformity may decrease. This may also make purging the cooling channel more difficult, which may be desirable in some embodiments because the weight of the electric aircraft may be minimized. However, it is undesirable for full turbulence to occur within the cooling channel, as the resulting pressure drop may make it difficult, impractical, or impossible to achieve high refrigerant flow rates (e.g., 100 or 200 liters / minute) which may be desirable in some embodiments. Finally, it is desirable to produce a dimple geometry that is cost-effective to manufacture while achieving these advantages.

[0151] Therefore, in some embodiments, the dimples 2210 may be arranged in a staggered pattern, as shown with respect to Figures 24A and 24B, to improve heat transfer and achieve optimal flow characteristics. This may offer significant advantages compared to dimples located in the center of the cooling channel. For example, centrally located dimples arranged periodically along the length of the channel may cause excessive flow separation at a low pitch (as seen in the dimple arrangement 2401 on the left side of Figure 24A) compared to a high pitch (as seen in the dimple arrangement 2402 on the right side of Figure 24A). By arranging the dimples on opposing sides of the channel in a staggered pattern (as seen in the dimple arrangement 2403 on the right side of Figure 24B), the refrigerant may follow a meandering path that does not cause excessive flow separation even with a dimple configuration that has a similar pitch to dimple arrangement 2401. For example, the dimples may be configured to generate a turbulent refrigerant flow, with each dimple introducing local turbulence into a separate laminar flow downstream of the dimple without completely changing the flow from a laminar state to a turbulent state. In some embodiments, the channel may be configured to generate a flow with a Reynolds number between 100 and 2000 (based on the channel's hydraulic diameter without dimples).

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

[0153] In some embodiments, the cooling channels 2222 located within the heat exchanger assembly 2200 may have a relatively small channel height compared to the channel width when viewed from the installation direction. For example, if the height of the cooling channel is, for example, about 1.5 mm, the channel width may be about 18.2 mm. In such cases, the manifold sections (e.g., manifold channels 2224 and 2226) may have a channel height of about 5 mm and a channel width of about 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 replenishment of refrigerant as needed during ground maintenance of the aircraft 100.

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

[0155] In this embodiment, the heat exchanger assembly 2200 may have symmetrical inlets and outlets 2240 and 2250 for receiving and returning the refrigerant, respectively. That is, the inlets and outlets 2240 and 2250 of the heat exchanger may be configured to receive or discharge the heat transfer fluid. Specifically, the symmetrical design of the heat exchanger assembly 2200 with respect to the inlets and outlets 2240 and 2250 of the heat exchanger can enable the implementation of a backflow strategy, as described above. For example, the direction of the refrigerant flow may be reversed as needed, for example, by switching between the inlets and outlets, by operating an external pump, or by 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, either before or after the first operating period.

[0156] Figure 22C shows an exemplary refrigerant flow path within a heat exchanger assembly 2200 for battery cooling in an HV cooling distribution subsystem, consistent with some embodiments of the present disclosure. As shown in Figure 22C, according to the embodiments of Figures 22A and 22B described above, the heat exchanger assembly 2200 may, for example, employ a “10P3S” or similar topology. Such a topology may include 10 cooling units (e.g., cooling unit 2230) connected in parallel, with each cooling unit having 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 distributed substantially uniformly among the 10 separate cooling unit paths, with each subdivision of the cooling fluid passing through all three cooling channels 2222a, 2222b, and 2222c of each of those cooling units 2230. In some embodiments, other related topologies such as 6P6S and 7P5S may be applied. Generally, the heat exchanger assembly 2220 may comprise a plurality of 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 provide inlets and outlets in the opposing manifold channels.

[0158] In some embodiments, the total volume of refrigerant in the heat exchanger assembly 2200 (i.e., excluding the volume of conduits in the fluid transport assembly) may be approximately 0.7 L. The 10P3S topology may 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 cross-flow design shown in Figures 22A-22C can reduce the thermal gradient that occurs when the temperature of the refrigerant rises. As shown in Figure 22C, in forward flow operation, after the heat exchanger assembly 2200 receives the refrigerant from the inlet (e.g., heat exchanger inlet / outlet 2240), the refrigerant flows into the first manifold section (e.g., the first manifold channel 2224) and is then distributed substantially evenly to a plurality of cooling units 2230. Each cooling unit 2230 may include, for example, three cooling channels connected in series (e.g., cooling channels 2222a, 2222b, and 2222c). The cooling channels within the heat exchanger assembly 2200 may be aligned parallel to each other. 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 in the figure, in each cooling unit 2230, the direction of flow may change when entering the next cooling channel. That is, the heat transfer fluid may flow in opposite directions in any two consecutive cooling channels within the same cooling unit 2230. For example, in forward flow operation, the refrigerant in cooling channels 2222a and 2222c flows to the right, as seen in Figure 22C, while the refrigerant in cooling channel 2222b flows to the left. After exiting the last cooling channel 2222c in cooling unit 2230, the refrigerant from the 10 cooling units flows into a second manifold section (e.g., a second manifold channel 2226) and exits through an outlet (e.g., a heat exchanger inlet / outlet 2250, as seen in Figure 22A).

[0160] In reverse flow operation, the heat exchanger assembly 2200 receives refrigerant from an inlet (e.g., heat exchanger inlet / outlet 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 the cooling unit 2230, the refrigerant from all 10 cooling units flows into a first manifold section (e.g., first manifold channel 2224) and exits through an outlet (e.g., heat exchanger inlet / outlet 2240 as seen in Figure 22A).

[0161] Figure 23A shows a magnified portion of the cooling channel 2300 within the heat exchanger assembly 2200 of Figure 22A, consistent with several embodiments of the present disclosure. Figure 23B shows a cross-sectional view along line AA of Figure 23A, consistent with several embodiments of the present disclosure. As shown in Figures 23A and 23B, the multiple 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 opposing lateral sides. In various embodiments, the dimensions of the cooling channel 2300 and the dimples 2310 can be designed to include specific implementations of the embodiments of the present disclosure, based on the actual needs for optimizing the overall heat transfer efficiency.

[0162] For example, design parameters that affect the refrigerant flow characteristics and, consequently, 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 vertices of two adjacent dimples 2310), the width 2312 of the dimple 2310, the height 2314 of the dimple 2310, the bottom angle of the dimple 2316 between the side wall of the cooling channel and the substantially flat, forward-sloping surface of the dimple, and the length 2318 of the vertex region of the dimple 2310. The optimal values ​​of such parameters may depend on the properties of the applied refrigerant, 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 this disclosure may offer several beneficial advantages and are suitable for cooling distribution systems in eVTOL aircraft. For example, the proposed designs reduce the mass of the heat exchanger assembly, as well as the required fluid mass in the cooling system. Designs with improved flatness also simplify the manufacturing process accordingly, improve overall thermal performance, and can meet important cell thermal requirements such as a low temperature gradient across each pack and a maximum temperature desirable for safety and performance reasons.

[0164] Embodiments of this disclosure may be further described by the following clauses. Clause 1. A cooling system for an electric aircraft, A plurality of heat exchanger assemblies corresponding to a plurality of battery packs included in the electric aircraft, each heat exchanger assembly comprising a first heat exchanger inlet and a second heat exchanger inlet configured to receive or discharge a heat transfer fluid, A fluid transport assembly coupled to the heat exchanger assembly, comprising a first main inlet and a second main inlet, and configured to circulate the heat transfer fluid in parallel with the heat exchanger assembly, wherein the first main inlet and the second main inlet are located at the same position on the electric aircraft, and the fluid transport assembly comprises, A cooling system further comprising a plurality of conduit segments configured to equilibrate the heat transfer fluid flowing into each of the heat exchanger assemblies. Clause 2. A cooling system according to Clause 1, wherein the plurality of conduit segments comprises a plurality of flow limiters configured to equalize the heat transfer fluid flowing into each of the heat exchanger assemblies. Clause 3. A cooling system as described in Clause 2, wherein each of the plurality of flow limiters comprises a sizing orifice configured to ensure substantially equal inlet pressures in each heat exchanger assembly in order to achieve a uniform flow distribution of the heat transfer fluid. Clause 4. A cooling system according to any one of Clauses 1 to 3, wherein each of the plurality of conduit segments includes a different inner diameter configured to equilibrate the heat transfer fluid flowing into each of the plurality of heat exchanger assemblies. Clause 5. A cooling system as described in any one of Clauses 1 to 4, wherein each heat exchanger assembly is: A heat exchanger plate, A first manifold channel coupled to the inlet and outlet of the first heat exchanger, A cooling system further comprising a heat exchanger plate, which includes a second manifold channel coupled to the second heat exchanger inlet / outlet. Clause 6. The cooling system described in Clause 5, A cooling system 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 turbulence of the heat transfer fluid through the cooling channel, each of the plurality of dimples having a symmetrical shape. Clause 7. A cooling system according to Clause 5 or 6, wherein the symmetrical shape comprises a first substantially flat and inclined surface viewed from a first flow direction of the heat transfer fluid and a second substantially flat and inclined surface viewed from a second flow direction of the heat transfer fluid opposite to the first flow direction. Clause 8. A cooling system as described in any one of Clauses 5 to 7, wherein the symmetrical shape includes one of a triangle and one of a trapezoid. Clause 9. A cooling system according to 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, the plurality of cooling units being coupled in parallel between the first manifold channel and the second manifold channel. Clause 10. A cooling system according to 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 and discharge the heat transfer fluid to the second heat exchanger inlet and discharge during a first operating period, and to receive the heat transfer fluid from the second heat exchanger inlet and discharge the heat transfer fluid to the first heat exchanger inlet and discharge during a second operating period before or after the first operating period. Clause 11. A cooling system according to any one of Clauses 5 to 10, wherein the heat exchanger plate comprises aluminum. Clause 12. A cooling system according to 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 opposing sides of the cooling channel. Clause 13. A cooling system according to Clause 12, wherein the adjacent dimples in the flow direction are recessed from the laterally opposing sides of the top or bottom surface of the cooling channel. Clause 14. A cooling system according to Clause 12, wherein the adjacent dimples in the flow direction are recessed from the laterally opposing side of the cooling channel. Clause 15. A cooling system as described in any one of Clauses 1 to 14, wherein the heat transfer fluid comprises ethylene glycol water (EGW), ethylene glycol, polyethylene glycol, water, or any combination thereof. Clause 16. A cooling system according to 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 connected to the first tube, A cooling system further comprising a second tube and a plurality of second hoses connected to the second tube. Clause 17. A cooling system according to Clause 16, wherein the first tube comprises a first main channel through which the first main inlet / outlet is connected, the plurality of first hoses comprises a first branch connected to the first main channel, the second tube comprises a second main channel through which the second main inlet / outlet is connected, and the plurality of second hoses comprises a second branch connected to the second main channel. Clause 18. A cooling system as described in Clause 17, wherein a plurality of flow limiters are connected to the first hose or the second hose. Clause 19. A cooling system according to any one of Clauses 16 to 18, wherein the first tube and the second tube are made of aluminum. Clause 20. A cooling system according to any one of Clauses 1 to 19, wherein the first main inlet and the second main inlet are located on the top side of the electric aircraft. Clause 21. A cooling system as described in any one of Clauses 1 to 20, wherein the fluid transport assembly is A first main road leading from the first main entrance to the first branch and the third branch, The system further comprises a second main road leading from the second main entrance to a second branch and a fourth branch, A cooling system in which the first trunk and the second trunk are substantially equal in length. Clause 22. A cooling system according to 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. Power supply system for an aircraft, A plurality of battery packs configured to be installed within the wing section of the aircraft, each battery pack configured to be connected to the electric engine of the aircraft to supply power, Multiple heat exchanger assemblies connected to the battery pack enclosures of the aforementioned multiple battery packs, The system comprises a fluid transport 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 comprises 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 having a plurality of dimples configured to provide turbulence for the heat transfer fluid, A power supply system in which each of the plurality of dimples includes a symmetrical shape having a first substantially flat and forward-sloping surface as viewed from a first flow direction of the heat transfer fluid and a second substantially flat and forward-sloping surface as viewed from a second flow direction of the heat transfer fluid opposite to the first flow direction. Clause 24. The power supply system described in Clause 23, wherein the heat exchanger plate is A first manifold channel coupled to the inlet and outlet of the first heat exchanger, A second manifold channel coupled to the inlet and outlet of the second heat exchanger, A power supply system comprising: 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. A power supply system as described in Clause 24, wherein each cooling channel comprises the plurality of dimples configured to provide the turbulence of the heat transfer fluid through the cooling channel. Clause 26. A power supply system according to Clause 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. Clause 27. A power supply system according to any one of Clauses 24 to 26, wherein each cooling channel comprises a plurality of dimples arranged in a staggered manner such that adjacent dimples in the first or second flow direction extend from opposing sides of the cooling channel. Clause 28. A power supply system as described in Clause 27, wherein the adjacent dimples in the first or second flow direction are recessed from the laterally opposing sides of the top or bottom surface of the cooling channel. Clause 29. A power supply system as described in Clause 27, wherein the adjacent dimples in the first or second flow direction are recessed from the laterally opposing side 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 and discharge the heat transfer fluid to the second heat exchanger inlet and discharge during a first operating period, and to receive the heat transfer fluid from the second heat exchanger inlet and discharge the heat transfer fluid to the first heat exchanger inlet and discharge during a second operating period before or after the first operating period. Clause 31. A power supply system according to any one of Clauses 24 to 30, wherein the heights of the first manifold channel and the second manifold channel are greater than the height of the cooling channel. Clause 32. A power supply system according to any one of Clauses 23 to 31, wherein the heat exchanger plate comprises aluminum. Article 33. A method for cooling a battery, A fluid transfer assembly circulates heat transfer fluid to multiple heat exchanger assemblies corresponding to multiple battery packs included in an electric aircraft, wherein each heat exchanger assembly comprises 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. The present invention provides turbulence of the heat transfer fluid through the cooling channels of the heat exchanger plate by means of a plurality of dimples, wherein the plurality of dimples in the cooling channels are arranged in a staggered pattern such that adjacent dimples in the flow direction of the cooling channels extend from opposing sides of the cooling channels, The operation of circulating the heat transfer fluid is as follows: During the first operating period, the heat transfer fluid is received from the first heat exchanger inlet and outlet, and the heat transfer fluid is discharged to the second heat exchanger inlet and outlet. A method comprising receiving the heat transfer fluid from the second heat exchanger inlet and outlet and discharging the heat transfer fluid to the first heat exchanger inlet and outlet during a second operating period. Clause 34. A method according to Clause 33, wherein the adjacent dimples in the flow direction are recessed from the laterally opposing sides of the top or bottom surface of the cooling channel. Clause 35. A method according to Clause 33, wherein the adjacent dimples in the flow direction are recessed from the laterally opposing side of the cooling channel. Clause 36. A method according to any one of Clauses 33 to 35, wherein each of the plurality of dimples includes a symmetrical shape having a first substantially flat and inclined surface as viewed from a first flow direction of the heat transfer fluid and a second substantially flat and inclined surface as viewed from a second flow direction of the heat transfer fluid opposite to the first flow direction. Article 37. The method described in any one of Articles 33 to 36, A method further comprising equalizing the heat transfer fluid flowing into the heat exchanger assembly by a plurality of flow limiters in the fluid transport assembly. Clause 38. The method described in Clause 37, wherein the heat transfer fluid flowing into the heat exchanger assembly is: A method comprising ensuring substantially equal inlet pressures in each heat exchanger assembly by using a sizing orifice for a uniform flow rate distribution of the heat transfer fluid. Clause 39. A method according to any one of Clauses 33 to 38, further comprising equalizing the heat transfer fluid flowing into the heat exchanger assembly by a plurality of conduit segments of different sizes in the fluid transport assembly. Clause 40. A method according to any one of Clauses 33 to 39, wherein the heat transfer fluid comprises ethylene glycol water (EGW), ethylene glycol, polyethylene glycol, water, or any combination thereof. Clause 41. A method according to any one of Clauses 33 to 40, wherein the dimple includes a triangular or trapezoidal shape. Article 42. Aircraft, A plurality of battery packs installed within the wing section of the aircraft, each battery pack being connected to the aircraft's electric engine to supply power, Multiple heat exchanger assemblies connected to the battery pack enclosures of the aforementioned multiple battery packs, The system comprises a fluid transport 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 comprises 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 having a plurality of dimples configured to provide turbulence for the heat transfer fluid, An aircraft in which each of the plurality of dimples includes a symmetrical shape having a first substantially flat and forward-sloping surface as viewed from a first flow direction of the heat transfer fluid and a second substantially flat and forward-sloping surface as viewed from a second flow direction of the heat transfer fluid opposite to the first flow direction. Article 43. Power supply system for an aircraft, A plurality of battery packs configured to be installed within the wing section of the aircraft, each battery pack configured to be connected to the electric engine of the aircraft to supply power, A plurality of heat exchanger assemblies corresponding to the plurality of battery packs included in an electric aircraft, each heat exchanger assembly comprising a first heat exchanger inlet and a second heat exchanger inlet configured to receive or discharge a heat transfer fluid, A fluid transport assembly coupled to the heat exchanger assembly, comprising a first main inlet and a second main inlet, and configured to circulate the heat transfer fluid in parallel with the heat exchanger assembly, wherein the first main inlet and the second main inlet are located at the same position on the electric aircraft, and the fluid transport assembly comprises, A power supply system further comprising a plurality of conduit segments configured to equilibrate the heat transfer fluid flowing into each of the heat exchanger assemblies. Article 44. Power supply system for an aircraft, A plurality of battery packs configured to be installed within the wing section of the aircraft, each battery pack configured to be connected to the electric engine of the aircraft to supply power, A plurality of heat exchanger assemblies connected to the battery pack enclosures of the plurality of battery packs, each heat exchanger assembly comprising a plurality of heat exchanger assemblies having a cooling channel, The fluid transport assembly is coupled to the plurality of heat exchanger assemblies and configured to circulate heat transfer fluid through the cooling channels within the heat exchanger assemblies, A power supply system in which each cooling channel has a plurality of dimples arranged in a staggered pattern such that adjacent dimples in the flow direction of the cooling channel extend from opposing sides of the cooling channel. Clause 45. A power supply system as described in Clause 44, wherein the adjacent dimples in the flow direction are recessed from the laterally opposing sides of the top or bottom surface of the cooling channel. Clause 46. A power supply system as described in Clause 44, wherein the adjacent dimples in the flow direction are recessed from the laterally opposing side of the cooling channel. Clause 47. A power supply system according to any one of Clauses 44 to 46, wherein each of the plurality of dimples includes a symmetrical shape having a substantially flat and forward-sloping surface as viewed from the direction of flow. Clause 48. A power supply system as described in Clause 47, wherein the dimple comprises a second substantially flat and forward-sloping surface when viewed from the opposite direction to the flow direction. Clause 49. A power supply system as described in 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 as described in any one of clauses 44 to 49. Article 51. Aircraft, An aircraft including a cooling system as described in any one of clauses 1 to 22. Article 52. Aircraft, An aircraft including a power supply system as described in any one of clauses 23 to 32. Article 53. Aircraft, A power supply system as described in any one of clauses 44 to 49, A cooling system as described in any one of clauses 1 to 22, An aircraft including one or more of the power supply systems described in any one of clauses 23 to 32.

[0165] The above description is provided for illustrative purposes only. It is not exhaustive and does not limit this disclosure to the exact form or embodiments disclosed. By considering the specifications and practices of the disclosed embodiments of the disclosure disclosed herein, modifications and adaptations of this disclosure will become apparent to those skilled in the art.

Claims

1. A cooling system for electric aircraft, A plurality of heat exchanger assemblies corresponding to a plurality of battery packs configured for the electric aircraft, each heat exchanger assembly comprising a first heat exchanger inlet and a second heat exchanger inlet configured to receive or discharge a heat transfer fluid, A fluid transport assembly coupled to the heat exchanger assembly, comprising a first main inlet and a second main inlet, configured to circulate the heat transfer fluid passing through the heat exchanger assembly in parallel, wherein the first main inlet and the second main inlet are configured to be located at the same position on the electric aircraft, and the fluid transport assembly comprises, The heat exchanger assembly further comprises a plurality of conduit segments configured to passively equilibrate the heat transfer fluid flowing into each of the heat exchanger assemblies, A cooling system comprising a heat exchanger plate, each heat exchanger assembly configured to receive the heat transfer fluid from a first heat exchanger inlet and discharge the heat transfer fluid to a second heat exchanger inlet during a first operating period, and to receive the heat transfer fluid from a second heat exchanger inlet and discharge the heat transfer fluid to a first heat exchanger inlet during a second operating period before or after the first operating period.

2. A cooling system according to claim 1, wherein the plurality of conduit segments comprises a plurality of fixed flow limiters configured to equalize the heat transfer fluid flowing into each of the heat exchanger assemblies.

3. A cooling system according to claim 2, wherein each of the plurality of fixed flow limiters comprises a sizing orifice configured to ensure equal inlet pressures in each heat exchanger assembly in order to achieve a uniform flow distribution of the heat transfer fluid.

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

5. A cooling system according to claim 1, wherein each heat exchanger plate is A first manifold channel coupled to the inlet and outlet of the first heat exchanger, A cooling system further comprising a second manifold channel coupled to the inlet and outlet of the second heat exchanger.

6. A cooling system according to claim 5, A cooling system 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 of the plurality of cooling channels comprising a plurality of dimples configured to provide turbulence of the heat transfer fluid through the cooling channel, each of the plurality of dimples having a symmetrical shape.

7. A cooling system according to claim 6, wherein the symmetrical shape comprises a first flat, forward-sloping surface viewed from a first flow direction of the heat transfer fluid and a second flat, forward-sloping surface viewed from a second flow direction of the heat transfer fluid opposite to the first flow direction.

8. A cooling system according to claim 6, wherein the symmetrical shape includes one of a triangle and a trapezoid.

9. A cooling system according to claim 6, 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.

10. A cooling system according to claim 1, wherein the heat exchanger plate comprises aluminum.

11. A cooling system according to claim 6, 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 opposing sides of the cooling channel.

12. A cooling system according to claim 11, wherein the adjacent dimples in the flow direction are recessed from laterally opposing sides of the top or bottom surface of the cooling channel.

13. A cooling system according to claim 11, wherein the adjacent dimples in the flow direction are recessed from the side surfaces of the cooling channel that are opposite to each other in the lateral direction.

14. A cooling system according to claim 1, wherein the heat transfer fluid comprises ethylene glycol water (EGW), ethylene glycol, polyethylene glycol, water, or any combination thereof.

15. A cooling system according to claim 1, wherein the plurality of conduit segments of the fluid transport assembly are A first tube and a plurality of first hoses connected to the first tube, A cooling system further comprising a second tube and a plurality of second hoses connected to the second tube.

16. A cooling system according to claim 15, wherein the first tube comprises a first main channel through which the first main inlet / outlet is connected, the plurality of first hoses comprises a first branch connected to the first main channel, the second tube comprises a second main channel through which the second main inlet / outlet is connected, and the plurality of second hoses comprises a second branch connected to the second main channel.

17. A cooling system according to claim 16, wherein a plurality of fixed flow limiters are connected to the first hose or the second hose.

18. A cooling system according to claim 15, wherein the first tube and the second tube are made of aluminum.

19. A cooling system according to any one of claims 1 to 18, wherein the first main inlet and the second main inlet are located on the top side of the electric aircraft.

20. A cooling system according to any one of claims 1 to 18, wherein the fluid transport assembly is A first main road leading from the first main entrance to the first branch and the third branch, The system further comprises a second main road leading from the second main entrance to a second branch and a fourth branch, The first and second trunks are of equal length in the cooling system.

21. A cooling system according to claim 20, wherein the first main inlet and the second main inlet are located closer to the first branch than to the second branch.

22. A method for cooling the battery, A fluid transfer assembly circulates heat transfer fluid to multiple heat exchanger assemblies corresponding to multiple battery packs included in an electric aircraft, wherein each heat exchanger assembly comprises 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. The present invention provides turbulence of the heat transfer fluid through the cooling channels of the heat exchanger plate by means of a plurality of dimples, wherein the plurality of dimples in the cooling channels are arranged in a staggered pattern such that adjacent dimples in the flow direction of the cooling channels extend from opposing sides of the cooling channels, The operation of circulating the heat transfer fluid is as follows: During the first operating period, the heat transfer fluid is received from the first heat exchanger inlet and outlet, and the heat transfer fluid is discharged to the second heat exchanger inlet and outlet. A method comprising receiving the heat transfer fluid from the second heat exchanger inlet and outlet and discharging the heat transfer fluid to the first heat exchanger inlet and outlet during a second operating period.

23. A method according to claim 22, wherein the adjacent dimples in the flow direction are recessed from laterally opposing sides of the top or bottom surface of the cooling channel.

24. A method according to claim 22, wherein the adjacent dimples in the flow direction are partially recessed from the laterally opposing sides of the top or bottom surface of the cooling channel.

25. A method according to claim 22, wherein the adjacent dimples in the flow direction are recessed from the laterally opposing side of the cooling channel.

26. A method according to claim 22, wherein each of the plurality of dimples includes a symmetrical shape having a first flat, forward-sloping surface as viewed from a first flow direction of the heat transfer fluid and a second flat, forward-sloping surface as viewed from a second flow direction of the heat transfer fluid opposite to the first flow direction.

27. A method according to claim 22, A method further comprising equalizing the heat transfer fluid flowing into the heat exchanger assembly by a plurality of flow limiters in the fluid transport assembly.

28. A method according to claim 27, wherein the heat transfer fluid flowing into the heat exchanger assembly is equal to A method comprising ensuring equal inlet pressure in each heat exchanger assembly by using a sizing orifice for a uniform flow rate distribution of the heat transfer fluid.

29. A method according to any one of claims 22 to 28, further comprising equalizing the heat transfer fluid flowing into the heat exchanger assembly by a plurality of conduit segments of different sizes in the fluid transport assembly.

30. A method according to any one of claims 22 to 28, wherein the heat transfer fluid comprises ethylene glycol water (EGW), ethylene glycol, polyethylene glycol, water, or any combination thereof.

31. A method according to any one of claims 22 to 28, wherein the dimple includes a triangular or trapezoidal shape.

32. A power system for aircraft, A plurality of battery packs configured to be installed within the wing section of the aircraft, each battery pack configured to be connected to the electric engine of the aircraft to supply power, A plurality of heat exchanger assemblies corresponding to the plurality of battery packs of the aircraft, each heat exchanger assembly comprising a first heat exchanger inlet and a second heat exchanger inlet configured to receive or discharge a heat transfer fluid, A fluid transport assembly coupled to the heat exchanger assembly and comprising a first main inlet and a second main inlet, configured to circulate the heat transfer fluid passing through the heat exchanger assembly in parallel, wherein the first main inlet and the second main inlet are configured to be located at the same position on the aircraft, and the fluid transport assembly comprises, The heat exchanger assembly further comprises a plurality of conduit segments configured to passively equilibrate the heat transfer fluid flowing into each of the heat exchanger assemblies, A power supply system comprising a heat exchanger plate, each heat exchanger assembly configured to receive the heat transfer fluid from a first heat exchanger inlet and discharge the heat transfer fluid to a second heat exchanger inlet and discharge it during a first operating period, and to receive the heat transfer fluid from a second heat exchanger inlet and discharge it to a first heat exchanger inlet and discharge it during a second operating period, either before or after the first operating period.

33. A power system for aircraft, A plurality of battery packs configured to be installed within the wing section of the aircraft, each battery pack configured to be connected to the electric engine of the aircraft to supply power, A plurality of heat exchanger assemblies connected to the battery pack enclosures of the plurality of battery packs, each heat exchanger assembly comprising a plurality of heat exchanger assemblies having a cooling channel, The fluid transport assembly is coupled to the plurality of heat exchanger assemblies and configured to circulate heat transfer fluid through the cooling channels within the heat exchanger assemblies, Each cooling channel comprises a plurality of dimples arranged in a staggered pattern such that adjacent dimples in the flow direction of the cooling channel extend from opposing sides of the cooling channel. A power supply system in which the adjacent dimples in the flow direction are partially recessed from the laterally opposing sides of the top or bottom surface of the cooling channel such that the first height of the cooling channel in the region without dimples is greater than the second height of the cooling channel in the region with dimples.

34. A power supply system according to claim 33, wherein the adjacent dimples in the flow direction are recessed from the side surfaces of the cooling channel facing each other laterally.

35. A power supply system according to claim 33, wherein each of the plurality of dimples includes a symmetrical shape having a flat, forward-sloping surface when viewed from the direction of flow.

36. A power supply system according to claim 35, wherein the dimple has a second flat, forward-sloping surface when viewed from the opposite direction to the flow direction.

37. A power supply system according to claim 33, wherein the dimples include a triangular or trapezoidal shape.

38. A power supply system according to any one of claims 33 to 37, wherein the fluid transport assembly is The first main entrance and, The second main entrance and The heat exchanger assembly comprises a plurality of conduit segments configured to passively equilibrate the heat transfer fluid flowing into each of the heat exchanger assemblies, The first main entrance and the second main entrance are power systems located at the same positions on the aircraft.

39. A power supply system according to claim 38, wherein the plurality of conduit segments comprises a plurality of fixed flow limiters configured to equalize the heat transfer fluid flowing into each of the heat exchanger assemblies.

40. A power supply system according to claim 39, wherein each of the plurality of fixed flow limiters comprises a sizing orifice configured to ensure equal inlet pressures in each heat exchanger assembly in order to achieve a uniform flow distribution of the heat transfer fluid.

41. A power supply system according to claim 38, wherein each of the plurality of conduit segments includes a different inner diameter configured to equilibrate the heat transfer fluid flowing into each of the plurality of heat exchanger assemblies.

42. The power supply system according to claim 33, wherein each heat exchanger assembly is A first manifold channel coupled to the inlet and outlet of the first heat exchanger, A second manifold channel coupled to the inlet and outlet of the second heat exchanger, A power supply system further equipped with a heat exchanger plate.

43. A power supply system according to claim 42, wherein the plurality of cooling channels are coupled between the first manifold channel and the second manifold channel.

44. A power supply system according to claim 42, 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.

45. A power supply system according to claim 42, wherein the heat exchanger plate is configured to receive the heat transfer fluid from the first heat exchanger inlet and discharge the heat transfer fluid to the second heat exchanger inlet and discharge during a first operating period, and to receive the heat transfer fluid from the second heat exchanger inlet and discharge the heat transfer fluid to the first heat exchanger inlet and discharge during a second operating period before or after the first operating period.

46. A power supply system according to claim 42, wherein the heat exchanger plate comprises aluminum.

47. A power supply system according to claim 33, wherein the heat transfer fluid comprises ethylene glycol water (EGW), ethylene glycol, polyethylene glycol, water, or any combination thereof.

48. The power supply system according to claim 38, wherein the plurality of conduit segments of the fluid transport assembly are A first tube and a plurality of first hoses connected to the first tube, A power supply system further comprising a second tube and a plurality of second hoses connected to the second tube.

49. A power supply system according to claim 48, wherein the first tube comprises a first trunk line through which the first main inlet / outlet is connected, the plurality of first hoses comprises a first branch connected to the first trunk line, the second tube comprises a second trunk line through which the second main inlet / outlet is connected, and the plurality of second hoses comprises a second branch connected to the second trunk line.

50. A power supply system according to claim 49, wherein a plurality of fixed flow limiters are connected to the first hose or the second hose.

51. A power supply system according to claim 48, wherein the first tube and the second tube are made of aluminum.

52. A power supply system according to claim 38, wherein the first main entrance and the second main entrance are located on the top side of the aircraft.

53. The power supply system according to claim 38, wherein the fluid transport assembly is A first main road leading from the first main entrance to the first branch and the third branch, The system further comprises a second main road leading from the second main entrance to a second branch and a fourth branch, The first and second trunk lines are of equal length in this power supply system.

54. A power supply system according to claim 53, wherein the first main inlet and the second main inlet are located closer to the first branch than to the second branch.

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